Whole document - Les Missions Scientifiques du Centre National d ...

PRO.LB.0.NT.003.ASC Issue. 6 rev. 3 Page: 3.1 

Chapter 4: Payload general design requirements 

CHANGE TRACEABILITY Chapter 4 

Here below are listed the changes between issue N-2 and issue N-1: 

N°§ PUID Change Status Reason of Change Change Reference 

§4.1.2.2 [PL - 4.1.2 -6 ] New in PUM.6.2.EJ.08 

§4.2.1.1 [PL - 4.2.1 - 2 a] Modified in STR cable added CIIS.4.1.JC.1_1 

§4.2.2.4 [PL - 4.2.2 - 5 a] Modified in Vertical handling PUM.6.1.CG.31_27 

§4.2.2.4 New in Nota added PUM.6.1.CG.31_27 

§4.2.2.4 New in Figure 4.2-3 added PUM.6.1.EJ.25 

§4.4.3.2 [PL - 4.4.3 - 7 a] Modified in AWG limitation PUM.6.1.EJ.15 

§4.6.1.5.2 Modified in X Co-coordinate modified PUM.6.1.EJ.33 

§4.7 [PL - 4.5.7 -1 ] Deleted in Replaced by PL-4.7-1 PUM.6.1.CG.31_23 

§4.7 [PL - 4.7 -1 ] New in New numbering+ additional 

sentence 

All right reserved. ALCATEL SPACE /CNES 

Passing and copying of this document, use and communication of its content is not permitted without prior written authorization. 

PUM.6.1.CG.31_23 

§4.7 [PL - 4.5.7 -2 ] Deleted in Replaced by PL-4.7-2 PUM.6.1.CG.31_23 

§4.7 [PL - 4.7 -2 ] New in New numbering PUM.6.1.CG.31_23


Here below are listed the changes from the previous issue N-1: 


§4.1.4.1 [PL - 4.1.4 - 2 a] Modified in To be documented in IDS/ICD PUM.6.2.EJ.08 

§4.1.4.1 [PL - 4.1.4 - 3 a] Modified in To be documented in IDS/ICD PUM.6.2.EJ.08 

§4.1.8.4 New in New Section: Storage requirements PUM.6.2.EJ.09 

§4.1.8.4 [PL - 4.1.8 -4 ] New in Storage requirements PUM.6.2.EJ.09 

§4.1.8.4 [PL - 4.1.8 -5 ] New in Storage requirements PUM.6.2.EJ.09 

§4.2.5.2 Modified in Safety factors for characterized 

materials added 



PUM.6.2.EJ.11 

§4.2.5.2 [PL - 4.2.5 - 4 a] Modified in Criteria added PUM.6.2.EJ.12 

§4.4.2.1 [PL - 4.4.2 -21 ] New in Sneak circuits and unintentional 

alctrical paths to be precluded 

§4.4.2.7.4 [PL - 4.4.2 -20 ] New in Connectors with electroexplosive 

devices 

§4.7 [PL - 4.7 -3 ] New in Warnings and precautions in PL AIT 

instructions 

PUM.6.2.EJ.30 

PUM.6.2.EJ.30 

PUM.6.2.EJ.14


TABLE OF CONTENTS 

4. CHAPTER 4: PAYLOAD GENERAL DESIGN REQUIREMENTS 9 

4.1 GENERAL DESIGN REQUIREMENTS 9 

4.1.1 INTERFACE CONTROL 9 

4.1.2 MATERIAL PROCESSES AND PARTS 10 

4.1.2.1 Parts and materials 10 

4.1.2.2 Magnetic materials 10 

4.1.2.3 Outgassing 10 

4.1.2.4 Threads and locking devices 11 

4.1.3 IDENTIFICATION AND PRODUCT MARKING 11 

4.1.4 MOUNTABILITY, INTERCHANGEABILITY AND ADJUSTMENT 11 

4.1.4.1 Hardware Accessibility 11 

4.1.4.2 Software Accessibility 11 

4.1.5 SAFETY 12 

4.1.6 CLEANLINESS 12 

4.1.7 AIT SUPPORT 12 

4.1.8 PREPARATION FOR STORAGE AND DELIVERY 13 

4.1.8.1 Retention of cleanliness 13 

4.1.8.2 Marking of the container 13 

4.1.8.3 Handling procedure 13 

4.1.8.4 Storage requirements 13 

4.2 MECHANICAL DESIGN REQUIREMENTS 14 

4.2.1 PAYLOAD PHYSICAL CHARACTERISTICS 14 

4.2.1.1 Mass 14 

4.2.1.2 Center of Gravity 14 

4.2.1.3 Moments of inertia 14 

4.2.1.4 Size 14 

4.2.2 PAYLOAD MOUNTING 15 

4.2.2.1 Method 15 

4.2.2.2 Grounding point 15 

4.2.2.3 Purging and venting interfaces 15 

4.2.2.4 Handling attach fittings/fixture 16 

4.2.3 ALIGNMENT 19 

4.2.4 CO-ALIGNMENT 20 

4.2.5 STRUCTURAL DESIGN 20 

4.2.5.1 Stiffness requirements 20 

4.2.5.2 Safety factors and safety margins 20 

4.2.5.3 Notching philosophy 21 

4.2.5.4 Structural mathematical models 21 

4.3 THERMAL DESIGN REQUIREMENTS 22 

4.3.1 DEFINITIONS 22 

4.3.1.1 Operational temperatures 22 

4.3.1.2 Acceptance temperatures 22 

4.3.1.3 Qualification temperatures 22 

4.3.1.4 Non operating temperatures 22 

4.3.1.5 Start-up temperatures 22 

4.3.1.6 Storage temperatures 22 


Passing and copying of this document, use and communication of its content is not permitted without prior written authorization.


4.3.2 THERMAL INTERFACES 23 

4.3.2.1 Thermal mathematical models and analysis 23 

4.4 ELECTRICAL DESIGN REQUIREMENTS 24 

4.4.1 SYSTEM GROUNDING 24 

4.4.1.1 General 24 

4.4.1.2 Structural grounding 28 

4.4.1.3 Thermal grounding 28 

4.4.1.4 Electrical bonding 29 

4.4.2 CABLING SHIELDING AND GROUNDING 30 

4.4.2.1 General 30 

4.4.2.2 Serial digital data acquisition, serial digital commands and low level commands grounding 30 

4.4.2.3 Digital relay acquisitions, and relay commands grounding 30 

4.4.2.4 Bi-level acquisitions grounding 31 

4.4.2.5 Thermistors acquisition and heaters commands grounding 31 

4.4.2.6 Analog signals grounding 31 

4.4.2.7 EED 31 

4.4.2.8 Shielding 33 

4.4.3 HARNESS REQUIREMENTS 34 

4.4.3.1 Pins assignment 34 

4.4.3.2 Harness design 34 

4.4.4 ISOLATION 36 

4.4.5 CONNECTORS TYPE AND KEYING 39 

4.5 COMMAND AND CONTROL DESIGN REQUIREMENTS 40 

4.5.1 GENERAL CONVENTIONS 40 

4.5.1.1 Word and byte convention 40 

4.5.1.2 Level 1 and 0 Conventions 40 

4.5.2 PROCESSOR TURN-ON TIME 41 

4.6 MATHEMATICAL MODELS INTERFACES REQUIREMENTS 42 

4.6.1 MECHANICAL MATHEMATICAL MODEL INTERFACES REQUIREMENTS 42 

4.6.1.1 General 42 

4.6.1.2 General Requirements 42 

4.6.1.3 Requirements for the dynamic models 47 

4.6.1.4 Requirements for correlated models 51 

4.6.1.5 Specific requirement for the payload 51 

4.6.2 THERMAL MODELS 52 

4.6.3 CAD MODELS 52 

4.7 SAFETY REQUIREMENTS 53 




LIST OF FIGURES 

Figure 4.2-0 : Ground stud configuration.............................................................................................................. 15 

Figure 4.2-1 : Satellite vertical handling procedure and payload exclusion areas (Calipso example)....................... 17 

Figure 4.2-3 : Satellite vertical handling................................................................................................................ 18 

Figure 4.2-2 : Satellite horizontal handling procedure (Calipso example)............................................................... 18 

Figure 4.4-1 : Grounding concepts ....................................................................................................................... 25 

Figure 4.4-2 : Symbols for grounding diagrams.................................................................................................... 26 

Figure 4.4-3 : Example of grounding diagram ...................................................................................................... 27 

Figure 4.4-4 : Common mode voltage.................................................................................................................. 37 

Figure 4.4-5: Signal interference isolation, in common mode ................................................................................ 38 




LIST OF TABLES 

Table 4.2-1: JY and JU safety factors for non pressurized item .............................................................................. 20 

Table 4.2-2: JY and JU safety factors for pressurized item ..................................................................................... 20 

Table 4.4-1:PROTEUS Bonding Requirement......................................................................................................... 29 

Table 4.5-1: Bit numbering inside a byte............................................................................................................... 40 

Table 4.6-1: Authorised NASTRAN cards .............................................................................................................. 47 

Table 4.6-2: Prohibited NASTRAN cards ............................................................................................................... 47 

Table 4.6-3: Numbering Range of the payload ..................................................................................................... 52 

Table 4.6-4: Co-ordinates of the payload-platform I/F nodes................................................................................ 52 




LIST OF CHANGE TRACEABILITY 

CHANGE TRACEABILITY Chapter 4 ........................................................................................................................ 1 

TABLE OF CONTENTS............................................................................................................................................ 3 

LIST OF FIGURES ................................................................................................................................................... 5 

LIST OF TABLES...................................................................................................................................................... 6 

LIST OF CHANGE TRACEABILITY ............................................................................................................................ 7 




LIST OF TBCs 

. 

Section Sentence Planned 

Resolution 

§4.2.2.2 The Payload Grounding Point shall be located as close as possible to the –Zs –Ys 

attachment foot and this location shall be identified in the payload ICD. This 

Grounding Point shall consist in a ground stud as shown in Figure 4.2-0 (TBC values 

are typical values which shall be defined by the Payload Supplier depending on 

payload design). Moreover, this Grounding Point shall be redundant (so 2 ground 

studs). 

§4.2.2.4 Nota: If the payload is non compliant with this exclusion area or if the Prime 

Contractor wants to handle the satellite by the payload, the payload handling attach 

fittings shall allow a vertical handling of the whole maximum equipped satellite. the 

whole maximum equipped satellite to be considered is the satellite maximum mass 

as indicated on Table 3.1-1 + 80 kg (TBC including all the non-flight harware (GSE) 

mounted on both the payload + platform + test PAF with associated separation 

device when handled). 

. 

List of TBDs 




4. CHAPTER 4: PAYLOAD GENERAL DESIGN REQUIREMENTS 

This chapter defines all the design requirements the payload shall comply in order to be compatible with the 

PROTEUS platform. It presents a generic specification concerning the general, mechanical, thermal, electrical design 

between the satellite bus and the payload. 

For the phase A of a PROTEUS based mission, the User is encouraged to contact either ALCATEL SPACE or CNES in 

order to consider every specification notified hereafter, check whether it is applicable to the studied payload and 

foresee the specific analysis necessary to adapt PROTEUS to the mission requirements. 

4.1 GENERAL DESIGN REQUIREMENTS 

4.1.1 INTERFACE CONTROL 

PL - 4.1.1 - 1 

The Payload Interface Control Document (see section 10.2) shall contain at least : 

• Payload Interface Data Sheet (Payload IDS framework with its filling rules are given in appendix) 

• Mechanical Interface Data Sheets (IDS) 

• Thermal IDS 

• Power data sheets : average power consumption, transient power demand and average power 

dissipation 

• List of connectors 

• Pin allocation data sheet 

• Elementary acquisitions and commands description sheet (discrete acquisition cycle) 

• Description of acquisition and command via serial lines 

• Description of acquisition and command via 1553 Bus 

• Telecommand data sheet (number of messages per unit, delay constraint between 2 consecutive TC) 

• Telemetry data sheet (number of messages per unit) 

• Miscellaneous sheet (all magnetic components, non flight hardware, location of purge valves and 

venting holes) 

• Drawings 

• Grounding scheme 

• Interfaces Description Drawings which completes IDS (volume, location of purge valves and venting 

holes, type and location of handling fixtures, definition of the conductive and radiative interface...) 

• Interfaces Description Documents for functional aspects. 

A copy of the Interface Data Sheet model will be provided by the Satellite Contractor in Excel software for PC, version 

5.0a, on a 3.5’’ floppy disk support. 




PL - 4.1.1 - 2 

This model shall be filled by the Payload Contractor and supplied to the Satellite Contractor in paper and 

software form in the same format. This format is described in appendix. 

PL - 4.1.1 - 3 

The definitions (mass, power...) given in appendix in the filling rules shall be applied at payload level. 

PL - 4.1.1 - 4 

Masses shall be established and recorded in the IDS and the record shall account for all mass status and 

mass dynamics attributable to deployable, consumable, moving or jettisonable materials or assemblies. 

PL - 4.1.1 - 5 

The external finishes of the payload (MLI, coatings, finishes...) shall be defined along with their optical 

properties at BOL and EOL in the payload ICD and/or IDS. 

PL - 4.1.1 - 6 

All the interfaces shall be defined and documented using the international system of units (metric, SI). 

4.1.2 MATERIAL PROCESSES AND PARTS 

4.1.2.1 Parts and materials 

PL - 4.1.2 - 1 

Material used at payload mechanical interface shall be compliant with Aluminium, steel and Titanium alloys. 

The use of pure Mercury, Cadmium and Zinc is prohibited. 

4.1.2.2 Magnetic materials 

PL - 4.1.2 - 2 

Non-magnetic materials shall be used for all components of payload except where magnetic materials are 

essential to the function of the unit. 

PL - 4.1.2 -6 

All magnetic components shall be clearly identified in an Interface Control Drawing. 

4.1.2.3 Outgassing 

PL - 4.1.2 - 3 

The Total Mass Loss (TML) of the payload shall be less than 1%. 

PL - 4.1.2 - 4 

The Collected Volatile Condensable Material (CVCM) of the payload shall be less than 0.1%. 




4.1.2.4 Threads and locking devices 

PL - 4.1.2 - 5 

Every bolted assembly (STA and connectors brackets) on the payload shall include a positive locking device 

(such as ONDUFLEX washers, for instance). 

4.1.3 IDENTIFICATION AND PRODUCT MARKING 

PL - 4.1.3 - 1 

The payload shall be permanently marked in French or English. The identification shall be visible when 

installed on the platform. The identification shall be visible with unaided eye from a 0.5 m distance. 

PL - 4.1.3 - 2 

The payload shall carry an identification with at least the following information : 

• Payload Assembly Name 

• Identification Part Number 

• Serial Number 

4.1.4 MOUNTABILITY, INTERCHANGEABILITY AND ADJUSTMENT 

PL - 4.1.4 - 1 

It shall be possible to mount and dismount several times (typically 5) the payload for integration constraints. 

4.1.4.1 Hardware Accessibility 

PL - 4.1.4 - 2 a 

The payload shall not require assembly or disassembly to perform mounting on or dismounting from the 

spacecraft. 

The payload design shall permit easy access to mounting bolts and to test points and components that may 

require adjustment. 

These points shall be identified in the ICD. 

PL - 4.1.4 - 3 a 

Non-flight hardware shall be clearly identified (red tags or marks) and easily accessible and removable. 

These non-flight hardware shall be identified in the instruments units ICDs. 

4.1.4.2 Software Accessibility 

PL - 4.1.4 - 4 

If a payload software is to be modified during AIT operations (for uploads or software configuration 

changes, for instance), this shall be easily feasible through test connectors and EGSEs, without dismounting 

anything. 

If performed, this operation shall be under the Payload Supplier responsibility. 




PL - 4.1.4 - 5 

The payload shall be designed such as the required adjustments (mechanical, electrical) are feasible at 

satellite level during AIT operations. 

4.1.5 SAFETY 

PL - 4.1.5 - 1 

Warnings and precautions relative to personnel and payload safety and hazards shall be specified in 

payload handling, assembly, and test instructions. 

PL - 4.1.5 - 2 

Payload and GSEs shall be compliant with launch pad and mission safety requirements (mission dependent). 

4.1.6 CLEANLINESS 

Hardware shall be designed, manufactured, assembled and handled in a manner to insure the highest practical level 

of cleanliness. 

PL - 4.1.6 - 1 

Suitable precautions shall be taken to insure freedom from debris within the hardware, and unaccessible 

areas where debris and foreign materials can become lodged, trapped, or hidden shall be avoided. 

PL - 4.1.6 - 2 

Hardware shall be designed so that malfunctions or inadvertent operations cannot be caused by exposure to 

conducting or non conducting debris or foreign materials floating in a gravity free state. 

PL - 4.1.6 - 3 

Electrical circuit shall be designed and fabricated to prevent unwanted current paths being produced by such 

debris. 

PL - 4.1.6 - 4 

All satellite related activities (after payload delivery) will be performed in class 100 000 clean rooms. The 

payload shall be compatible with this class. 

4.1.7 AIT SUPPORT 

PL - 4.1.7 - 1 

The Payload Supplier shall provide all the necessary tooling, equipment and working media for the payload 

assembly and test at system level. 




4.1.8 PREPARATION FOR STORAGE AND DELIVERY 

4.1.8.1 Retention of cleanliness 

PL - 4.1.8 - 1 

The payload shall be sealed for retention of cleanliness using precleaned bags as port closures and shall be 

retained by pressure sensitive tape applied over the bags. The payload shall be double bagged in antistatic 

polyethylene or polyamid film (100 micron total thickness minimum) and shall then be packed properly 

according to commercial practice in a manner which will provide adequate protection against hazards 

encountered during transportation, handling and/or storage. 

4.1.8.2 Marking of the container 

PL - 4.1.8 - 2 

The container for the payload shall be labelled, tagged or marked to permit detailed identification of the 

content of the container. 

4.1.8.3 Handling procedure 

PL - 4.1.8 - 3 

The payload handling procedure shall be delivered with the payload container and shall be accessible 

without opening it, in order to allow payload incoming inspection after delivery at Satellite Contractor 

Facilities. 

4.1.8.4 Storage requirements 

PL - 4.1.8 -4 

Special storage conditions and constraints, if any, shall be listed by the Payload Supplier. When integrated 

on the satellite, the payload may be stored in clean room environment up to 8 months. 

PL - 4.1.8 -5 

End of storage operations shall be minimized. 




4.2 MECHANICAL DESIGN REQUIREMENTS 

4.2.1 PAYLOAD PHYSICAL CHARACTERISTICS 

4.2.1.1 Mass 

PL - 4.2.1 - 1 

The payload mass shall include the total hardware that is intended to fly. 

PL - 4.2.1 - 2 a 

The payload mass shall be presented as follows : 

• Equipped payload mass, including the STA, H02 & H03 connectors brackets and STR cables mass, 

• mounting hardware such as screws, washers, bonding straps or equivalent, interface fillers when 

delivered. 

4.2.1.2 Center of Gravity 

PL - 4.2.1 - 3 

The Center of Gravity (CoG) shall be identified related to the Payload Reference Frame (shown section 1.4). 

4.2.1.3 Moments of inertia 

4.2.1.4 Size 

PL - 4.2.1 - 4 

Moments of inertia shall be identified related to the Payload reference Frame (show section 1.4). 

PL - 4.2.1 - 5 

The volume allocated to the payload includes the total hardware that is intended to fly. 

PL - 4.2.1 - 6 

The nominal external dimensions of the payload shall be expressed in millimeters. 

An overstepping of the maximum dimensions toward the allocated dimensions will induce a formal volume change 

notice. 

PL - 4.2.1 - 7 

The external envelope dimensions of the delivered hardware (excluding thermal blankets) shall not exceed 

the dimensions specified in the Interface Control Drawing by more than 1.25 mm unless specifically 

authorized. 




4.2.2 PAYLOAD MOUNTING 

4.2.2.1 Method 

PL - 4.2.2 - 1 

Payload mounting shall be accomplished by M8 bolts, passing through interface pods flanges, as described 

in section 3.1.4. 

Payload mounting bolts shall be provided by the Payload Contractor. 

4.2.2.2 Grounding point 

PL - 4.2.2 - 2 

The Payload Grounding Point shall be located as close as possible to the –Zs –Ys attachment foot and this 

location shall be identified in the payload ICD. This Grounding Point shall consist in a ground stud as shown 

in Figure 4.2-0 (TBC values are typical values which shall be defined by the Payload Supplier depending on 

payload design). Moreover, this Grounding Point shall be redundant (so 2 ground studs). 

(TBC) 

(TBC) 

Figure 4.2-0 : Ground stud configuration 

4.2.2.3 Purging and venting interfaces 

PL - 4.2.2 - 3 



(TBC) 

(TBC) 

The payload shall be vented to accommodate the specified barometric pressure rates of change for both 

decreasing and increasing pressure but shall avoid any pollution problem during tests or handling on 

ground.


4.2.2.4 Handling attach fittings/fixture 

PL - 4.2.2 - 5 a 

These handling attach fittings shall allow a vertical handling of the equipped Payload. Handling required 

environments (covering the environment encountered at ALCATEL SPACE facilities) are given in section 

5.11.2.1.3 and required safety factors in section 4.2.5.2. 

The payload mass to be considered shall include all the non-flight hardware mounted on the payload when 

handled. 

In addition, in case of hazardous system handling, a fail-safe analysis (loss of one handling point) shall be 

performed with environments required in section 5.11.2.1.3 and safety factors given in section 4.2.5.2. 

For safety reasons, this sizing shall be approved by ALCATEL SPACE. 

PL - 4.2.2 - 6 

deleted 

The Satellite handling will be directly performed through platform dedicated handling MGSE. Schematics are 

provided in figure 4.2-1 and 4.2-2 to illustrate the foreseen handling procedures. 

PL - 4.2.2 - 7 

deleted 

PL - 4.2.2 - 8 

The payload shall comply with the exclusion areas indicated on the figure 4.2-1 for vertical handling MGSE 

(including slings). No additional exclusion area is required for the satellite horizontal handling 

configuration(see PL-3.1.3-3 for the payload allowed volume). 

Nota: If the payload is non compliant with this exclusion area or if the Prime Contractor wants to handle the satellite 

by the payload, the payload handling attach fittings shall allow a vertical handling of the whole maximum equipped 

satellite. the whole maximum equipped satellite to be considered is the satellite maximum mass as indicated on 

Table 3.1-1 + 80 kg (TBC including all the non-flight harware (GSE) mounted on both the payload + platform + 

test PAF with associated separation device when handled). 

Handling required environments (covering the environment encountered at ALCATEL SPACE facilities) are given in 

Section 5.11.2.1.3 and required safety factors in Section 4.2.5.2. 

In addition, in case of hazardous system handling, a fail-safe analysis (loss of one handling point) shall be 

performed with environment required in Section 5.11.2.1.3 and safety factors given in Section 4.2.5.2. 

For safety reason , this sizing shall be approved by ALCATEL SPACE. 

In this case, tets at payload level shall be performed in order to demonstrate the possibility of handling the whole 

satellite at ALCATEL facilities. These tests shall be performed before delivery and with additional masses in order to 

represent the maximal satellite mass. 

The maximum load encountered during nominal handling shall be tested on the flight hardware. 

The maximum load encountered during degraded case (fail-safe for instance) shall be tested on a representative 

sample. 

For safety resons, related test reports shall be provided with the payload in its acceptance. 




Figure 4.2-1 : Satellite vertical handling procedure and payload exclusion areas (Calipso example) 

.. 




Figure 4.2-3 : Satellite vertical handling 

Figure 4.2-2 : Satellite horizontal handling procedure (Calipso example) 




4.2.3 ALIGNMENT 

The integration / alignment of the Payload on the Platform is a Satellite Contractor responsibility but it is the Payload 

Supplier responsibility to provide a payload design which is compatible with the required alignment accuracy. The 

nominal alignment measurements during satellite AIT aim at verifying the stability of the payload master reference 

cube between the beginning of the AIT mechanical and thermal environment tests and the end of these tests. In 

addition, position of this cube will be measured with respect to the STR reference cube in order to determine the 

relative position of the payload boresight with respect to the STR boresight. No other optical cubes are nominally 

controlled during satellite AIT. 

PL - 4.2.3 - 1 

Consequently, a description of payload specific alignment method or needs during satellite AIT shall be 

provided by the Payload Supplier for Satellite Contractor approval before Satellite phase C. 

This shall include as a minimum a definition of the payload adjustment range, a detailed description of the 

hardware used for that purpose, the reference cube and its field of view. 

It shall be noticed that, at satellite level, all alignment measurements will be performed in vertical configuration 

(Satellite +Xs axis in vertical position). This leads to some constraints at payload level expressed in PL - 4.2.3 - 3. 

PL - 4.2.3 - 2 

For each feed/antenna reflector assembly, a specific device (palmer equipped struts, as far as possible) 

allowing verification of the relative geometry shall be provided to the Satellite Contractor by the Payload 

Supplier, if geometry verification after satellite testing (mechanical, thermal, EMC) is required at System level. 

PL - 4.2.3 – 3 

The payload master reference optical cube shall fulfil the following requirements : 

• This cube shall be duplicated (one nominal and one redundant) and the 2 cubes shall not be located on 

the same area (full redundancy rules) 

• These 2 cubes shall be directly accessible with the satellite in vertical position that is to say by 2 

perpendicular horizontal lines of sight (+Ys and +Zs for instance or any combination remaining in 

+Ys+Zs). Moreover, the previous line of sight shall be free of any interference, 

• They shall be located on a stable area (moreover, they shall be preferably located on the +/- Ys faces of 

the payload) 

• Their minimum size shall be 20 mm x 20 mm x 20 mm, 

• They shall not be dismounted during payload or satellite test campaign. 

Information about alignment references, accuracy and adjustments form part of the PL ADP. Any other payload 

optical cube, if measured during satellite AIT, shall be compliant with PL-4.2.3-3 except for the redundancy aspect. 




4.2.4 CO-ALIGNMENT 

N.A 

4.2.5 STRUCTURAL DESIGN 

4.2.5.1 Stiffness requirements 

PL - 4.2.5 - 1 

The first mode frequencies of the payload required in the chapter 3 shall be achieved with the assumption of 

a hard-mounted interface on an infinitely rigid interface. 

4.2.5.2 Safety factors and safety margins 

PL - 4.2.5 - 2 

For non pressurized item, the following safety factors for yield sizing (JY) and for ultimate sizing (JU), with 

respect to the qualification loads, shall be applied : 

Type of hardware JY (yield) JU (ultimate) 

Metal flight hardware 1.25 1.56 

Metal inserts and joints (flight hardware) NA 2.0 

with characterized materials(*) NA 1.25 

Composite flight hardware NA 1.56 


Composite inserts and joints (flight hardware) NA 2.0 


Ground handling 1.5 2.5 

(*) : the applicable safety factors may be reduced for pieces of hardware made of materials characterized through an 

appropriate number of tests, yielding to a better knowledge of the admissible stress ("A" or "B" type). 

PL - 4.2.5 - 7 

Table 4.2-1: JY and JU safety factors for non pressurized item 

For pressurized item, the following safety factors for yield sizing (JY) and for ultimate sizing (JU), with respect 

to the Maximum Operating Design Pressure (defined as the highest pressure occurring from maximum relief 

pressure, maximum regulator pressure, maximum temperature or transient pressure excursions), shall be 

applied. 

In addition, pressure vessels design and verification shall comply with the requirements specified in Safety 

Regulations requirements. 

Type of hardware JY (yield) JU (ultimate) 

Pressure vessels 1.25 1.56 

Pressurized components 1.5 2.5 

Table 4.2-2: JY and JU safety factors for pressurized item 




PL - 4.2.5 - 3 

The following definitions shall be applied: 

Qualification loads = 1.25 x Flight limit loads except for ground handling where: 

- qualification loads = 2 x Flight limit loads for nominal analysis and 

- qualification loads = 1.5 x limit loads for fail safe analysis (where 1.5 corresponds to the dynamic 

factor due to the loss of one handling point). 

Sizing loads = JY (or JU) x Qualification loads 

The safety margin is defined as follows : 

where : 

σ admissible 

S. 

M. 

= −1 

σ 

calculated 

the admissible stress is the yield (respectively ultimate) stress when estimating the safety margin wrt yield 

(respectively ultimate). 

the admissible stress for single points failure pieces of hardware shall be "A" type values (99% probability with 

a 95% confidence level), 

the admissible stress for redundant pieces of hardware shall be "B" type values (90% probability with a 95% 

confidence level), 

the calculated stress is the qualification stress times the yield (respectively ultimate) safety factor. 

PL - 4.2.5 - 4 a 

All safety margins shall be positive: 

• local buckling criterion: there shall be no buckling under qualification loads 

4.2.5.3 Notching philosophy 

PL - 4.2.5 - 5 

Notching at payload level is allowed during sine vibration test in order not to exceed, at the 

platform/payload interface, the Quasi-Static equivalent loads (resultant forces and moments at the 

geometrical centre of the interface points) after Satellite Contractor agreement 

Notching at payload units level is forbidden. 

PL - 4.2.5 - 6 

For notching during random vibration tests, the Payload Supplier shall contact either ALCATEL SPACE or 

CNES. 

4.2.5.4 Structural mathematical models 

The requirements for the quality of mathematical models, numbering ranges and interface point locations are given 

in section 4.6.1. 




4.3 THERMAL DESIGN REQUIREMENTS 

4.3.1 DEFINITIONS 

For information, the temperatures definition used for the PROTEUS platform are given hereafter. 

4.3.1.1 Operational temperatures 

The minimum and maximum operational temperatures TOPMIN and TOPMAX are the extreme temperatures a 

payload shall withstand during its specified lifetime for its various operational modes. 

4.3.1.2 Acceptance temperatures 

Acceptance temperature limits shall be deduced from operational temperature limits by an extension of 5°C : 

TAMIN = TOPMIN - 5 °C 

TAMAX = TOPMAX + 5 °C 

4.3.1.3 Qualification temperatures 

Qualification temperature limits shall be deduced from operational temperature limits by an extension of 10°C : 

TQMIN = TOPMIN - 10 °C 

TQMAX = TOPMAX + 10 °C 

4.3.1.4 Non operating temperatures 

Minimal and maximal non operating temperatures TNOPMIN and TNOPMAX are the extreme temperatures a 

payload shall withstand when it is OFF during specific satellite modes or during ground phase up a few days (for 

example transport). 

4.3.1.5 Start-up temperatures 

Minimal and maximal start up temperatures TSUMIN and TSUMAX are the extreme temperatures a payload shall be 

able to be turned ON, possibly without fulfilling all its performance requirements (for example a «cold start » after 

modes transition, when the satellite stayed in Safe mode for a long time before coming back to normal mode). 

4.3.1.6 Storage temperatures 

Minimal and maximal temperatures TSTOMIN and TSTOMAX are the extreme temperatures a payload shall 

withstand during storage phase up to several months. 




4.3.2 THERMAL INTERFACES 

4.3.2.1 Thermal mathematical models and analysis 

PL - 4.3.2 - 1 

A thermal analysis is required at payload level covering the most thermally critical operating modes and 

including transient cases where relevant. 

In a standard approach, no Payload reduced mathematical model is required. Nonetheless, the platform thermal 

control design has to take into account radiative coupling with the payload. 

PL - 4.3.2 – 2 

So, the Payload Supplier shall provide an external radiative geometrical model, in flight (before payload 

PDR) and satellite-level test configurations (six months before test) in the Payload Interface Control 

Document. 

This model is an external representation of the Payload including for each main surface : 

• Thermo-optic characteristics 

• Worst temperatures assumption (cold and hot). 




4.4 ELECTRICAL DESIGN REQUIREMENTS 

4.4.1 SYSTEM GROUNDING 

4.4.1.1 General 

PL - 4.4.1 - 1 

The satellite structure shall not be used as a current carrying conductor. 

PL - 4.4.1 - 2 

Shields shall not be used for signal returns except for RF signals. 

PL - 4.4.1 - 3 

An overall zero volt and grounding diagram shall be provided in the ICD/IDS for assessing the functional 

and electromagnetic compatibilities. This diagram shall indicate any AC or DC loop, the type of isolation 

used, any impedance coupling between zero volt and structure, and the type of connection between 

secondary 0 V and mechanical ground (if any). 

PL - 4.4.1 - 4 

The Payload Supplier shall provide a grounding diagram based on the following concepts. 




PROTEUS Power 

Strap 

Battery 

supply 

PL - 4.4.1 - 5 

PCE 

Converter 

Primary 0V 

Secondary 0V 

Chassis ground 

DHU 

Unit 1 

Unit 2 

Unit 3 

Unit 4 

Unit 5 

Unit 6 

ZVS1 

ZVS2 

ZVS1 

Figure 4.4-1 : Grounding concepts 

The following rules and symbols shall be used to draw grounding diagrams 

Solution 1 

Solution 2 

Solution 3 



For RF parts of 

equipment only 

Solution 4

.. 


i 

: Chassis ground 

: Ground 

: Secondary 0V n°i 

: Primary 0V 

: Bonding stud 

: Twisted pair 

: Twisted shielded pair 

: Coxial cable 

: DC/DC converter (isolated) 

: Signal transformer 



M 

T° 

: Single ended amplifier (transmitter) 

: Differential amplifier (transmitter) 

: Single ended amplifier (receiver) 

: Differential amplifier (receiver) 

: Optocoupler 

: Motor 

: Thermistor 

: Heater 

: Metallic housing grounded via mountin 

: Metallic housing grounded via foil strip 

Figure 4.4-2 : Symbols for grounding diagrams


Converter 

Primary 0V 

Secondary 0V 

Chassis ground 

ZVS1 

ZVS2 

Figure 4.4-3 : Example of grounding diagram 




4.4.1.2 Structural grounding 

PL - 4.4.1 - 6 

All structural members of the satellite and payload chassis and enclosures shall be designed to provide 

electrical conductivity across all mechanical joints, except where DC isolation is required for maximum 

electrical reliability. Conductive surface protection coatings such as Iridite, Alodine, or plating shall be used 

at all joints. The DC resistance across fixed joints shall not exceed 2.5 mOhm. 

PL - 4.4.1 - 7 

Redundant bonding straps shall be employed across joints where direct metal-to-metal contact cannot be 

assured. The DC resistance of these straps shall not exceed 10 mOhm. 

4.4.1.3 Thermal grounding 

PL - 4.4.1 - 8 

The conductive surfaces of all metal or metallic coated thermal components, such as heat shields and 

metallized layers of thermal blankets (that shall include one conductive layer) shall be electrically grounded 

to the satellite structure with a DC resistance lower than 10 mOhm. The number of bonding points per sheet 

of MLI shall be compliant with the following rules: 

• Sheets of 0.5m2 max: two points, at corners of the longest diagonal, as a minimum, 

• Sheets of 1 m 2 max: four points, at each corner, as a minimum, 

• Sheets greater than 1 m 2 : one bonding point at diagonal corners and intermediate points along outer 

sheet edges to ensure bonding areas not to exceed 1 m 2 . 

PL - 4.4.1 - 9 

In addition, the resistance between two bonding points in a MLI shall be lower than 80 Ω. 

PL - 4.4.1 - 10 

The use of non conductively coated insulators shall be minimized. 




4.4.1.4 Electrical bonding 

PL - 4.4.1 - 11 

Electrical bonding shall be in accordance with Table 4.4-1 with the following additions : 

a. The exterior case, including connectors and all metallic external covers shall be electrically bonded, 

directly or indirectly to chassis ground with a resistance of no greater than 2.5 milliohm per bond except for 

composite components. 

b. For composite components, the DC resistance per bond shall be no greater than 100 Ohm. 

c. The mounting surface shall be such that it may be electrically bonded to the structure upon which it is to be 

mounted at installation in the spacecraft. 

d. All internal mechanical assemblies shall be electrically bonded directly or indirectly to the base plate. 

e. Gimbaled, hinged, or jointed interfaces shall be bonded by means of redundant grounding straps. 

Bonded configuration max DC resistance 

(Ohm) 

RF boxes to Panel Ground Reference (PGR) 0.010 

Non-RF boxes to PGR 0.020 

Electrical boxes on graphite panels (if any) to PGR 0.050 

Across hinges (antenna deployed booms & solar array) 0.100 

Units, optical heads to PGR 0.020 

Harness shields to PGR 0.020 

Antenna to PGR 300.0 

Thermal blankets ground to Single Ground Point (SGP) 0.010 

Mechanical equipment to SGP 1.0 

Thermal blanket to multiple grounding tab to tab 0.010 

Thermal shields (thrusters) to structure 1.0 

Panel Ground Reference to SGP 0.10 

Table 4.4-1:PROTEUS Bonding Requirement 




4.4.2 CABLING SHIELDING AND GROUNDING 


PL - 4.4.2 -21 

The design shall preclude sneak circuits and unintentional electrical paths 

PL - 4.4.2 - 1 

The primary electrical power distribution system will have the power negative grounded to the spacecraft 

structure at a single ground point (SGP). 

PL - 4.4.2 - 2 

The electronic boxes supporting structure shall be designed with a panel ground reference (PGR). The PGR 

shall consist of ground studs, or inserts for ground straps, to be connected between the panel and the 

adjacent panels. 

The DC resistance between PGR and panel structure shall be lower than 10 mΩ. 

PL - 4.4.2 - 3 

Secondary power return lines shall be connected to the equipment structure in a single point. Exceptions are 

RF communication equipments and electrical units with operating frequency > 10 MHz where the secondary 

return can be connected to the structure with a lot of points. 

PL - 4.4.2 - 4 

Command signal wiring: 

In general, the wiring for the command signals shall be implemented using 26 gauge twisted pair wire from 

the branch module. Command signal with rise times < 200 µs which are routed through harness paths 

common to signal wires with susceptibility thresholds less than 10 V and less than 10 ms pulse response 

time, shall be shielded. 

PL - 4.4.2 - 5 

Each power line shall be electrically isolated with a dedicated return. 

4.4.2.2 Serial digital data acquisition, serial digital commands and low level commands grounding 

PL - 4.4.2 - 6 

Serial digital data acquisition and command signals shall be carried on shielded twisted pair wires and shall 

use structure as signal reference. Receiver shall be isolated from the primary ground ; emitter shall be 

ground referenced. 

Low level commands shall also be carried on shielded twisted pairs. 

4.4.2.3 Digital relay acquisitions, and relay commands grounding 

PL - 4.4.2 - 7 

Digital relay acquisitions and relay commands shall be completely electrically isolated with dedicated return. 




4.4.2.4 Bi-level acquisitions grounding 

PL - 4.4.2 - 8 

Bi-level acquisitions shall use secondary power. 

4.4.2.5 Thermistors acquisition and heaters commands grounding 

PL - 4.4.2 - 9 

Thermistors/heaters lines shall be electrically isolated. Thermistors acquisitions shall use twisted shielded pair 

wires. Heaters commands shall use twisted pairs. 

4.4.2.6 Analog signals grounding 

4.4.2.7 EED 

PL - 4.4.2 - 10 

Each analog acquisition line shall use a dedicated return which will be grounded at user end. 

Analog signals at interfaces shall be arranged to allow the use of twisted shielded wire. 

Exception for high accuracy, analog acquisition which shall not be grounded. 

4.4.2.7.1 General EED Wiring. 

PL - 4.4.2 - 11 

All EED wiring circuits shall use double shielded twisted pair wires. The return side of the circuits shall be 

grounded at the power supply; exceptions shall be submitted to Satellite Contractor for approval. The shields 

shall be grounded at the connector backshell at all connectors. 

4.4.2.7.2 EED Circuit Shields 

PL - 4.4.2 - 12 

Firing circuit shields shall provide a minimum of 20 dB attenuation from 30 kHz to 18 GHz. All firing circuit 

bundles shall use a double shielding configuration that has zero aperture from the power control unit to the 

electroexplosive devices (EED). The inner shield on these harnesses shall be the regular flat braided shield of 

the cables, which provides a minimum coverage of 90%. The outer shield shall be an overall shield to 

provide complete coverage from end to end. There shall be no gaps or discontinuities in the shielding, 

including the terminations at the back faces of the connectors. Electrical continuity and isolation of the inner 

and outer electroexplosive circuit shields shall be maintained. 

4.4.2.7.3 EED Cabling 

PL - 4.4.2 - 13 

Bundles shall be manufactured such that several electroexplosive device circuits are contained in a common 

shielded bundle. Splices within the bundles are forbidden. When breaking of a circuit is required, a mating 

connector pair shall be provided. All bundles shall be routed as close to the conductive metal ground plane 

of the platform/payload structure as feasible, with provision for tie-downs a maximum of 15 mm apart. 




4.4.2.7.4 EED Circuit Connectors 

PL - 4.4.2 - 14 

Connectors used in the electroexplosive bundles shall be of the circular MIL-C-26482 Series 2 type with 

conductive nickel slated metal shell bodies. These shall be self-locking and compatible with the mating 

equipment interface connectors. There shall be only one wire per pin, and in no case shall a connector pin 

be used as a terminal or a tie-point for multiple connections. 

PL - 4.4.2 -20 

All connectors used with the electroexplosive devices shall: 

• be approved by the procuring activity, 

• have a stainless steel shell or suitable electrically conductive finish, 

• complete the shell-to-shell connection before the pins connect, 

• provide for 360° shield continuity. 

There shall be only one wire per pin, and in no case shall a connector pin be used as a terminal or a tiepoint 

for multiple connections. 

The source circuits shall terminate in a connector with socket contacts. 

Connectors shall be selected such that they are not subject to demating when exposed to the maximum 

anticipated environment. 

Connectors that twist and lock into position are preferred. 

4.4.2.7.5 EED Circuit Redundant Wiring 

PL - 4.4.2 - 15 

Redundant EED circuits shall be wired and routed in separate wire bundles where required. Separation of 

wire bundles shall be maintained to the maximum extent possible, including the use of separate connectors if 

feasible. 

4.4.2.7.6 EED Harness Electrical Bonding 

PL - 4.4.2 - 16 

The EED harness hardware shall be bonded to the spacecraft local panel ground reference through the 

mating equipment chassis. Each connector and shield termination shall be assembled (mated) and tested to 

insure a maximum resistance of 10 mΩ (between connector or shield and PGR). 

4.4.2.7.7 EED Harness identification. 

PL - 4.4.2 - 17 

Each EED harness shall be positively identified by part number and serial number. Identifying information 

may be attach directly to the wiring harness by a sleeve attach to the harness. Other forms of identification 

such as mylar nameplates, metal nameplates, metal stampings, vibropeening, acid, electrical or 

mechanically etched, embossed, forged, brazed, cast or molded methods of manufacture shall not be used. 




4.4.2.7.8 EED Harness Records 

PL - 4.4.2 - 18 

Each EED wiring harness assembly shall have inspection and test records maintained by appropriate number 

and with a connector mate log maintained for all connectors from initial assembly and test throughout unit 

and satellite integration and acceptance test lifetime. 

4.4.2.8 Shielding 

PL - 4.4.2 - 19 

The general shielding guideline shall be such that each line function is evaluated to determine if it is possible 

to cable the line as an unshielded wire. 

The shielding guidelines are as follows : 

a. All telemetry lines shall be shielded individually or in functional groups. 

b. Unit interfaces which interconnect with sensitive or susceptible circuitry or are proximate to sensitive or 

susceptible circuitry shall be shielded. 

c. All command lines may use shielded wiring (except the digital command lines which may use unshielded 

wires). 

d. Regulated power lines shall use shielded wire. 

e. Shields shall not carry currents (except RF). 

f. Shields shall be jacketed to provide isolation from ground and chassis except at designated points. 

g. Shields shall provide a minimum of 90 percent coverage (e.g. tinned copper braid and woven copper). 

h. All pyrotechnics lines shall be double shielded. Firing circuits shielding shall provide a minimum of 20 dB 

attenuation from 30 kHz to 18 GHz. All firing circuit harnesses shall use a double shielding configuration 

that has zero aperture from the DHU to the electroexplosive devices (EED). There shall be no gaps or 

discontinuities in the shielding, including the terminations at the back faces of the connectors. Electrical 

continuity and isolation of EED circuit shields shall be maintained. All electrical cables may be fabricated 

such that several EED circuits are contained in a common shield cable bundle. There shall be no splices 

within the cable bundles. 

i. Each end every cable or waveguide going through the Payload shall have shield bonded to the payload 

structure over 360 deg. 




4.4.3 HARNESS REQUIREMENTS 

PL - 4.4.3 - 1 

Circuits having incompatible electromagnetic interference characteristics shall be segregated in cabling and 

connectors to the maximum extent possible to minimize interference coupling. 

Separation is necessary for the following circuits categories : 

DC power and command circuits 

Digital signals (0 - 5 V) 

Analog signals (0 - 5 V) 

Electro Explosive Devices 

Radio Frequency lines 

Wires carrying proprietary data 

MIL-STD-1553B bus 

4.4.3.1 Pins assignment 

PL - 4.4.3 - 2 

If two or more circuit categories must share a connector, pin assignments shall be made to provide a 

maximum of isolation in the connector and facilitate separation of the wiring external to the connector. A 

minimum of two pins separation shall be used. 

4.4.3.2 Harness design 

PL - 4.4.3 - 3 

Signal control interface harnesses, in general shall be constructed using twisted shielded wires. Signal return 

lines shall be shielded. However, some pulse commands and relay driver lines may not be shielded in order 

to save weight on the satellite. In this case, EMI analysis shall be performed to ensure EMC/ESD requirement 

compliance. 

PL - 4.4.3 - 4 

Neither the structure nor any cable shield shall be used to carry power bus return. This will minimize 

common mode noise input to the units. 

PL - 4.4.3 - 5 

For sensitive and critical functions, another shield shall be added that is continuous from the backshells of 

each of the associated unit connectors. 

PL - 4.4.3 - 6 

All shields shall be terminated to chassis external to the unit enclosure. Where external cables penetrate the 

enclosure of the satellite main body, they shall be terminated to the structure externally. 




PL - 4.4.3 - 7 a 

The following general rule on wiring shall be applied unless special approval has been granted: 

a. wiring shall be sized to provide a maximum voltage drop for power lines of 240 mV between source 

and user. For telemetry lines, this value can be as high as 1.0 V depending on the telemetry system 

parameter. 

b. wiring size and shield shall be : 

• No 20, 22 and 24 AWG twisted shielded pairs for secondary power distribution. If flexible wiring is 

utilized, it shall be EMI controlled, and in accordance with MIL-P-50884B. 

• No 26 AWG single through nine conductors twisted shielded wire for control and monitor. 

PL - 4.4.3 - 8 

For explosive parts, all circuits shall use twisted double shielded pair wires. The return side of the circuits shall 

be grounded at the power supply. Shield shall be grounded at both ends of the harness. 

PL - 4.4.3 - 9 

When feasible, redundant wiring shall be routed in separate wire bundles. Separation of wire bundles shall 

be maintained to the maximum extent possible, including the use of separate connectors, if necessary. 

PL - 4.4.3 - 10 

Spare wires shall not be provided in wiring harness terminating in removable crimp-contact connectors. 

PL - 4.4.3 - 11 

EMI controls on printed flexible wiring includes shielding and guard conductors. A circuit pattern may have 

shields on one or both sides. Additional shielding may be used on circuit edges if necessary. Top and bottom 

shielding may be added as solid conductive material connected and tied electrically. Insulation layers 

(covercoats) are normally used as outside layer. Guard conductors are effective in reducing adjacent trace 

coupling (crosstalk). 

PL - 4.4.3 - 12 

Every cable submitted to the external environment (i.e external to the Payload Instrument Module) shall be 

overshielded. 




4.4.4 ISOLATION 

PL - 4.4.4 - 1 

Onboard power, supplied by inverters, converters and transformer isolated power supplies, shall be defined 

as secondary power and shall be referenced to the mechanical ground at one location, at the secondary 

power source, via the shortest possible low impedance path. For those types of equipments, secondary 

power return is connected to the mechanical ground. In this case, the DC resistance between secondary 

power return line and mechanical ground shall be less than 2.5 mΩ. 

Remote sensors, pressure sensors, magnetometers, or assemblies without internal power supply may be 

exempt from the above requirement and secondary power return is isolated from the mechanical ground. 

PL - 4.4.4 - 2 

Primary power : 

All the users shall maintain an electrical isolation of at least 1 MΩ shunted by not more than 50 nF between: 

• primary power positive and chassis, 

• primary power return and chassis, 

• primary power return and secondary power return. 

PL - 4.4.4 - 3 

Secondary power: 

except secondary single point referenced, all the sources and loads shall maintain an electrical isolation of at 

least 1 MΩ shunted by not more than 50 nF between: 

• secondary power positive and chassis, 

• secondary power return and chassis, 

• primary power return and secondary power return. 

At no time the satellite will impose more than 1.5 V DC and 1 V peak to peak, from 15 kHz to 180 KHz, falling to 

0.2 V at 15 MHz, between the primary return and secondary return. 




2 

1 

0 

PL - 4.4.4 - 4 

V pp 

Common mode voltage 

1.00E+04 1.00E+05 1.00E+06 1.00E+07 1.00E+08 

Frequency (Hz) 

Figure 4.4-4 : Common mode voltage 

Differential interface circuits between instrument units shall be designed to maintain a common mode 

isolation as described on the Figure 4.4-5. 




10 

Impedance in KOhm 

Signal interface isolation 

in common mode 

0 

1.00E+01 1.00E+02 1.00E+03 1.00E+04 1.00E+05 1.00E+06 1.00E+07 1.00E+ 


Figure 4.4-5: Signal interference isolation, in common mode 




4.4.5 CONNECTORS TYPE AND KEYING 

PL - 4.4.5 - 1 

Use of micro-D connectors is not allowed. 

PL - 4.4.5 - 2 

The MIL-STD-1553B bus connectors shall be dedicated (no sharing of connectors with any other signal) and 

segregated (one connector for nominal bus and one for redundant bus) on each unit using this bus. 

PL - 4.4.5 - 3 

Deleted 

PL - 4.4.5 - 4 

The payload shall employ connector keying, where required, to prevent accidental mismating of connectors. 

The harness mating connectors shall be configured to properly maintain this keying requirement The harness 

shall be designed to interface with the mating connectors of the spacecraft electrical units with provision for 

unit and harness serviceability after assembly. Access shall be provided which supports safe and proper 

mating and demating of all connectors after spacecraft integration. 

PL - 4.4.5 - 5 

Electrical connectors shall be electrically bonded to the metallic case in which they are installed to provide 

electrical resistance of less than 2.5 mOhm. Except for cases otherwise approved by the satellite contractor, 

the connector housing shall be bonded to the chassis via a strap with a resistance of less than 10 mOhm. 




4.5 COMMAND AND CONTROL DESIGN REQUIREMENTS 

4.5.1 GENERAL CONVENTIONS 

PL - 4.5.1 - 1 

The following conventions shall be used at payload level. 

4.5.1.1 Word and byte convention 

A word is composed of 16 bits. 

A byte is composed of 8 bits. 

The numbering of the bits inside a byte shall be as follows : 

Integer bit B0 B1 B2 B3 B4 B5 B6 B7 

decimal value (MSB) 

(LSB) 

1 0 0 0 0 0 0 0 1 

128 1 0 0 0 0 0 0 0 

255 1 1 1 1 1 1 1 1 

Table 4.5-1: Bit numbering inside a byte 

Note : there is an equivalent convention for a word, yielding to B0 as MSB and B15 as LSB. 

4.5.1.2 Level 1 and 0 Conventions 

Convention for direct commands : 

TC 1 level shall reflect the ON or ENABLE command to the concerned circuit: 

active level of a relay, 

closed contact of a switch. 

TC 0 level shall reflect the OFF or DISABLE command to the concerned circuit: 

quiescent level of a relay, 

open contact of a switch. 

Convention for serial commands: 

when applicable, logic one voltage (TC 1) level shall reflect the ON or ENABLE command to the concerned 

circuit, MSB shall be transmitted first. 

Convention for direct acquisitions: 

TM 1 level shall reflect the ON or ENABLE status of the concerned circuit: 

closed contact of a relay, 

logic one of a status. 

TM 0 level shall reflect the OFF or DISABLE status of the concerned circuit: 

open contact of a relay, 

logic zero of a status. 

Convention for serial acquisitions: 




when applicable, logic one voltage (TM 1) level shall reflect the ON or ENABLE status of the concerned circuit, 

MSB shall be transmitted first. 

4.5.2 PROCESSOR TURN-ON TIME 

PL - 4.5.2 - 1 

Full reset, start up, and initialization maximum duration shall be provided in their IDS for payload with flight 

electronic processor inside. 

There shall be no polling of these units until they are declared operational by the Ground Segment. 




4.6 MATHEMATICAL MODELS INTERFACES REQUIREMENTS 

PL - 4.6 - 1 

The delivered mathematical models shall comply with the following requirements (section 4.6.1 to 4.6.3). 

4.6.1 MECHANICAL MATHEMATICAL MODEL INTERFACES REQUIREMENTS 


This section presents the general requirements for the delivered mathematical models which will be mounted on 

PROTEUS platform Finite Element Model (FEM). 

The system structural analysis will be performed with the finite element code MSC/NASTRAN. 

Therefore, all delivered mathematical models are required in NASTRAN format. These models will be used to 

perform system analyses. 

The Payload Supplier must provide two kinds of model: a physical model and a reduced model (condensed or modal 

model). 

preliminary physical and reduced models due date: at the beginning of the satellite phase B 

detailed physical and reduced models due date: at the beginning of the satellite phase C/D 

one correlated reduced model due date: after payload qualification test 

The models shall be in accordance with the following items defined in the next paragraphs: 

utilisation of versions compatible with version 70 of the NASTRAN Code, 

the basic axis system and the payload system, 

the rules of modelisation, 

the interface nodes: co-ordinates, boundaries conditions, number of degrees of freedom (d.o.f.) and 

identification number of the GRID cards, 

the loaded nodes: number of d.o.f. and identification number of the GRID cards, 

the conditioning of the matrices, 

the form of the delivery. 

For information, the products of inertia are defined with the following sign convention : 

Ixy = - x y dm ; Iyz = - y z dm ; Ixz = - x z dm 

4.6.1.2 General Requirements 

4.6.1.2.1 Axis systems and payload system 

The basic axis system of the payload model shall be the satellite reference frame (show section 1.4). 

The unit system is the International System (meter, kilogram, second, radian). 

Local axis systems are prohibited for the definition of the co-ordinates and the degrees of freedom (displacements) of 

all the conserved nodes (loaded and interface ones). 

All local axis systems must be defined wrt the basic one, and their number limited to around 5. 

4.6.1.2.2 Rigid bodies 

Any rigid body or rigid element connecting interface nodes between them is prohibited. 




If an interface node is connected by a rigid element to non interface nodes, the d.o.f of the interface node must be 

the independent (or non constrained) d.o.f. 

4.6.1.2.3 Data and capabilities requested 

The physical finite elements model shall contain all the data necessary to perform eventually : 

a static analysis 

a modal analysis (eigen modes calculation) 

a sine response analysis 

a thermoelastic analysis (only Coefficient of Thermal Expansion (CTE) and reference temperature). 

Concerning the thermoelastic analysis, the finite element model shall be such that it does not introduce stresses 

higher than 1 MPa in the payload primary structure under the following loading case : 

payload with isostatic boundary conditions 

coefficient of thermoelastic expansion set to 20 10 -6 °C -1 (only for this test) on all parts of the model 

homogeneous increase of temperature of +100°C applied to the whole model. 

One of the major condition necessary to fulfil this requirement is that there will be no rigid body with length > 0 

connecting 2 nodes of the payload primary structure. 

4.6.1.2.4 Masses representation 

The masses representation (choice between concentrated masses and distributed masses) is to be defined by the 

supplier in order to fit as well as possible with the actual payload masses distribution. However, the meaning of the 

masses representation cards will be explained by comments cards. 

In case of distributed mass, the following data are requested : 

structural mass value (kg/m, kg/m 2 or kg/m 3 depending on the element) 

non structural mass values (kg/m, kg/m 2 or kg/m 3 depending on the element) 

total mass value (kg/m, kg/m2 or kg/m3 depending on the element). 

The model rigid mass along the 3 axes must have the same value.(use of CMASS2 elements could generate 

problems). 

4.6.1.2.5 Results of the non condensed model (physical model) 

The following data are requested: 

description of the model with plots of the mesh showing clearly the numbering of the most important nodes 

(conserved, loaded, interface, ...) and elements 

results of the eigen mode analysis performed under free-free boundary conditions (frequencies of the 6 rigid 

modes + frequencies of the 3 first elastic modes) 

results of the eigen mode analysis performed with the specified boundaries conditions (frequencies, effective 

masses and inertia, participation factors, plots of mode shapes) for the significant modes 

masses, inertia, centre of gravity of the model compared with the data of the real mass breakdown 

results of the test defined § 4.6.1.2.10 and § 4.6.1.2.3, allowing to state on the acceptability of the F.E.M. with 

regard to thermoelastic analysis requirements. 




4.6.1.2.6 Results of the condensed model 

The results considered in this paragraph correspond to the condensed model with the specified boundaries 

conditions. For the free-free condensed model, only the values of the 9 first frequencies are required. 

The following data are requested: 

same results and plots that for the uncondensed model, excepted thermoelastic test 

In addition, a table of comparison to demonstrate the agreement between the condensed and the uncondensed 

models, containing for both models: the frequencies, the effective masses and inertia, the participation factors up to 

150 Hz at least or until the residual masses and inertia are less than 20% (or higher if test sequence at sub-system 

level foresees larger frequency band). 

In the frequencies range defined above, the requested representativity for the condensed model modal characteristics 

with regard to the physical model ones shall be: 

±5% on frequencies 

±15% on effective masses and inertia. 

4.6.1.2.7 Data about the condensation 

"Conserved nodes (or d.o.f.)" means the loaded nodes (or d.o.f.) plus the interface ones. 

The following data and deliveries are requested whatever form of the delivered model: 

the partitioning vector to expanse the condensed matrices (size nc x nc) to the physical matrices (size nt x nt) 

under the form of a column vector (DMI NASTRAN vector) where: 

nc is the number of conserved d.o.f. 

nt is the number of conserved nodes multiplied by 6. 

the conserved GRID cards package: each conserved node shall be present on a GRID card. All the d.o.f. of 

the conserved nodes which do not appear in the matrices are to be permanently constrained to zero directly 

on the GRID card. 

if the interface nodes are loaded by any mass or inertia, that shall be clearly indicated, 

the ASET 1 cards package 

No resequencing process must be used for the creation of the condensed matrices. It is required to use the following 

NASTRAN parameter: 

PARAM,NEWSEQ,-1 

4.6.1.2.8 Plotel cards package 

Plotel elements connecting the conserved nodes are required to plot the undeformed and deformed structures with 

sufficient representativity. 

The identification number of the Plotel cards will be taken between the same limits that for the conserved nodes for 

the payload. 

To make more representative the plots of the substructure, some nodes could be especially conserved, but with the 

6 d.o.f. blocked as they would be used only for the figures. They must fulfil all the requirements of the conserved 

nodes. 

4.6.1.2.9 Check of the delivered condensed model 

The payload supplier shall verify that the stiffness matrix is well conditioned by the herebelow tests and shall show the 

results of these tests. 




A modal analysis has to be performed using large mass and NASTRAN SUPORT card in order to extract rigid modes, 

elastic modes and effective masses. The strain energy of the rigid modes and the conditioning parameter ε have to 

be low: 

STRAIN ENERGY < 10-3 J (Joule) 

ε < 10 -8 

The previous modal analysis with large mass can be replaced by a constraint check performed when the condensed 

model is in free-free configuration. The test to be performed is to calculate the strain energy as defined below for 

each rigid mode Φ R: 

where: 

S.E. = ½ .[ Φ R ] T .[K].[ Φ R ] 

• [ΦR ] is a vector for one of the six rigid modes 

• [ ΦR ] T • [K] 

the transposed vector of [ΦR ], 

is the model stiffness matrix 

A unit rigid body displacement is applied on the whole structure on the 6 DOF (3 translations and 3 rotations). 

The strain energy computed for each of these rigid body motions shall be: 

< 10 -3 J 

This last test shall be performed without SUPORT card. 

This Strain Energy Check is used to identify constrained or grounding problems in a FEM model and ensure that the 

model is free-free. 

This test can be performed using NASTRAN DMAP rigid body checks, or by a specific NASTRAN DMAP. 

A free-free modal analysis has to be performed without large mass, without SUPORT card and without rigid interface 

in order to extract the six rigid modes. The frequency of these modes divided by the first elastic mode shall be : 

< 10 -4 

Moreover, the 3 first elastic free-free frequencies will be provided. 

All these tests must be performed on the condensed matrices or on the physical model re-read on the delivered tape. 

All the models not in accordance with these tests will be rejected. 

4.6.1.2.10 Check of the delivered full model 

Idem § 4.6.1.2.9 

4.6.1.2.11 Magnetic tape characteristics 

The package defined here above will be provided on one of the following magnetic data storage: 

Streamer cartridge 150 Mbytes , SGI , UNIX , tar format 

DAT 4mm (60, 90 or 120 m length) SGI , UNIX , tar format 

Floppy disk 3 «1/2 DOS formatted ASCII code 

The command used for the creation of the tape archive is to be delivered with the tape delivery. 

If data are compressed, the uncompress software must be provided on the magnetic tape . 

4.6.1.2.12 Associated documentation 

The technical note delivered with the tape shall include all the items mentioned in chapter 2: 




the results of the non condensed model (§ 4.6.1.2.5), 

the results of the condensed model (§ 4.6.1.2.6), 

the results of the tests about the matrices conditioning (§ 4.6.1.2.9), 

plots of each substructure showing clearly the numbering of the nodes and of the elements of the physical 

model, 

plots performed with the PLOTEL cards showing the conserved nodes, 

a scheme showing the various local axis systems w.r.t. the basic one, 

a description of each substructure and of the way to modelise, 

an explanation of the modelisation hypotheses and of the equivalent representations, 

the detailed mass breakdown of the model compared with the real one and indications of: 

the representation of the masses: concentrated or distributed, 

the considered offsets and rigid bodies between the masses and the structure, 

a table summarising the main structural characteristics of each substructure 

a summary of the tape contents. 




4.6.1.3 Requirements for the dynamic models 

4.6.1.3.1 Physical models 

4.6.1.3.1.1 General 

The authorised NASTRAN elements are provide in the table hereunder. 

Item connectivity card property card material card 

1D elements PROD, PBAR, PBEAM MAT1 

2D elements CTRIA3, CQUAD4 PSHELL MAT1, MAT2, MAT8 

3D elements CPENTA, CTETRA, CHEXA PSOLID MAT1 

MAT9 

Masses CONM1, CONM2, CMASS2 - - 

Local stiffness and 

CELAS1, CELAS2 PELAS 

connection 

- 

Rigid element & constraint RBAR, RBE2, MPC, RBE3 - - 

Miscellaneous PLOTEL - - 

NASTRAN parameter PARAM AUTOSPC YES - - 

Table 4.6-1: Authorised NASTRAN cards 

The following NASTRAN cards are to be prohibited: 

NASTRAN prohibited cards 

PARAM BAILOUT 

PARAM K6ROT 

NASTRAN parameters * 

PARAM MAXRATIO 

PARAM EPZERO 

NASTRAN parameters ** PARAM WTMASS 

NASTRAN cards CQUAD8, CQUADR, CTRIAR CTRIA6, EGRID 

Table 4.6-2: Prohibited NASTRAN cards 

• * parameters affecting model conditioning 

• **parameters affecting the other models during FEM assembly 

In case of necessity to use other cards than the authorised elements, the supplier will have to ask for the agreement 

of ALCATEL SPACE. 

Remarks concerning FEM rules: 

It is requested to use elements RBE2 with zero length and MPC with zero length (to simplify the use for 

thermoelastic analyses) 

The use of MPC card is forbidden to link interface nodes between substructures. The interface shall be performed 

with CELAS or zero length RBE2. 

The interface nodes do not have to be dependent nodes of rigid bodies 

The interface nodes do not have to be linked together by a rigid body 

Only RBE3 with simply supported independent nodes will be allowed (independent nodes: DOF 123 (456 

forbidden) , reference node: No restriction on dependent DOF) 

4.6.1.3.1.2 Size limitation 




The physical models are limited to 3000 nodes and 3000 elements. 

4.6.1.3.1.3 List of the data to be supplied 

The model will be supply on magnetic tape with the characteristics given on § 4.6.1.2.11. 

The following items will be provided on the tape: 

the ASET1 cards, 

the CORD2 cards, 

the complete Bulk Data Deck gathered in one file. 

Preferably the BDD will be built in order that all the cards defining the same substructure will be gathered together in 

the file. Enough comments will be added to make easy the understanding of the file. 

4.6.1.3.2 Condensed physical models 

4.6.1.3.2.1 General 

The nodal points and degrees of freedom will be defined as follows: 

the degrees of freedom will be related to nodal points (6 dof maximum per nodal point ordered as follows: T x, 

T y, T z, R x, R y, R z - T for translation and R for rotation), 

nodal points will be supplied in ascending numerical order, 

nodal points co-ordinates will be supplied according to the satellite reference axes system (see §4.6.1.2.1). 

Local co-ordinates system are not acceptable. 

nodal points must be kept at the location of the accelerometers foreseen for the sine test vibrations 

nodal points must be chosen in order to plot the deformed shapes of the structure with a sufficient 

representativity 

These rules will determine the numbering of the rows and columns of the mass, stiffness and damping (if supplied) 

matrices. 

4.6.1.3.2.2 Size limitation 

The maximum size of the stiffness, mass and damping (if supplied) matrices including the interface dof is 500 x 500. 




the ASET1 cards package, 

the DMI partitioning vector, 

the conserved GRID cards, 

the PLOTEL cards, 

the stiffness, mass and damping (if supplied) matrices in NASTRAN format OUTPUT4 option BCD non sparse, 

with D23.16 format for version 68 and following. 




4.6.1.3.3 Modal models 

4.6.1.3.3.1 General 

The dynamic behaviour of the structure is described by the reduced stiffness and mass matrices, relative to the elastic 

cantilevered modes and rigid body modes of the structure. 

The motion of the structure is represented as a superposition of the rigid body and elastic cantilevered motions. 

The rigid body motion is represented by the six rigid body modes shapes referenced to unity at the structure interface. 

The elastic motion is represented by the elastic cantilevered structure interface modes shapes. 

Thus: 

where: 

( ) 

X 

= 

I 

Φ 

Φ 

 

0 

e R 

( p xn) ( p x6m) 

( 6mxn) ( 6mx6m) q 

 

X 

 

 

 

n is the number of elastic modes 

p is the number of dof of the source model matrices 

Φ R 

Φ e 

are the rigid body modes shapes 

are the elastic cantilevered modes shapes 

(normalized such that Φ e t MΦe=I* ) 

X is the motion of the structure degrees of freedom 

Using the above formulation, the modal equations of motion are: 

where: 

GEN 

* At for transposed A and I for identity matrix. 

or in partitionned formulation: 

where: 

M I 

K I 

µ 

 

t 

M el 

i i 

GEN 

cantilevered modal co - ordinates 

payload interface motion 

( M ) Q 

+ ( C ) Q 

+ ( K ) Q = F 

q 

Q = 

 

 

 

X i 



GEN 

GEN 

2 

M el 

 

Q 

 

2µξω 

0 

+ Q 

µω 0 0 

+ 

 

Q = 

M 

 

I 0 K I 0 K I FI 

 

 

 

() 1 

is the condensed stiffness matrix at interface 

is the condensed mass matrix at interface


B I 

M eI 

t ( ) e e MΦ Φ = 

is the condensed damping matrix at interface 

is the elastic coupling mass matrix 

µ is the generalized mass matrix (normalized to unity) 

( µξω ) ( 2ξω 

) 

2 = is the generalized damping matrix 

2 2 ( ) ( ω ) 

µω = is the generalized stiffness matrix 

( FI ) 

are the structure interface loads. 

The size of MGEN, CGEN and KGEN matrices is N rows by N columns where N is the number of elastic modes increased 

of the six interface rigid body degrees of freedom. These matrices must be diagonal and elastic modes with no 

modal mass are forbbiden. 

4.6.1.3.3.2 Restitution matrices 

For modal data delivery, displacement restitution matrix is necessary to provide structure internal responses. This 

matrix provide analytical relationship between the internal responses and the modal generalized parameters. This 

matrix must be a subset of the transformation matrix (see equation (1)) used to reduce the mass and stiffness 

matrices. 

The equation is: 

The restitution matrix must containt at least: 

( X ) = DTMQ 

Nodes at the location of the accelerometers foreseen for the sine test vibrations 

Nodes allowing to plot the deformed shapes of the structure with a sufficient representativity 

Any other nodes considered as important by the supplier. 

4.6.1.3.3.3 Matrices size and output requirement limitations 

The maximum size of the stiffness, mass and damping (if supplied) matrices including the interface dof is 500 x 500. 

The number of restitution parameters must be less or equal to the number of dof allowed for physical condensed 

model. 

The delivered modal model will contain internal nodes representative to the main parts of the subsystem in order to 

allow the exploitation of dynamic responses inside of modal model (Notching possibilities). 

For modal model, only the modes having their effective mass (or inertia) greater than 0.5 % of the total subsystem 

mass will have to be retained in model delivery. 




the list of dummy nodal points (see remark below), 




the list of dummy degrees of freedom (see remark below), 

the ASET1 cards package, 

the DMI partitioning vector, 

the restitution GRID cards, 

the PLOTEL cards, 

the stiffness, mass and damping (if supplied) matrices in NASTRAN format OUTPUT4 option BCD non sparse 

with D23.16 format for version 68 and following, 

the displacement restitution matrix format OUTPUT4. 

Remark: 

To allow provision of modal model in the same way as condensed physical model, definition of dummy nodal points 

and dummy degrees of freedom is necessary. These points and degrees will be defined as follows: 

n nodal points (with 1 dof) associated to the n elastic modes (numbered in the range defined in § 5.1). For 

each nodal point, the co-ordinates are (1., 0., 0.). 

m nodal points (with 6 dof) associated to each structure interface point. Their number have to be greater than 

n and must be chosen in the range given in § 4.6.1.5.1. Their co-ordinates are defined in § 4.6.1.5.2. 

The such defined number of degrees of freedom will be the same as the reduced mass and stiffness matrices size 

(N). 

4.6.1.4 Requirements for correlated models 

4.6.1.4.1 Purpose 

The purpose of the correlated models is to perform the System dynamic analyses to prepare the satellite system sine 

tests and to confirm the its behaviour in flight by means of a transient response and/or a Coupled Analysis with the 

launch vehicle. 

The models must be representative of the last definition of the hardware and of the tests results. For the correlation 

with the tests the goals are: 

< ± 5 % on the frequencies 

for the significant modes and mainly for the first ones. 

For this delivery: 

the physical F.E.M. are requested, 

a comparison between the test results and the test predictions of the correlated model has to be provided. 

4.6.1.4.2 Comparison between predictions and tests 

The supplier of the payload shall provide the comparison of the frequencies and amplifications measured during the 

low level runs and the ones predicted by the correlated model for : 

each main mode, 

the point with the highest response in the correlated F.E.M., 

instrumented point with the highest response during the test of each substructure, 

4.6.1.5 Specific requirement for the payload 

This chapter presents the requirements concerning the numbering range and the co-ordinates of the interfaces with 

the platform. 




4.6.1.5.1 Numbering range of the payload 

Grids, elements 

rigid bodies, MPC 

Allowed Numbering Range 

Properties Materials Co-ordinate 



system 

Payload 50001-90000 50001-90000 50001-90000 50001-90000 

Table 4.6-3: Numbering Range of the payload 

The physical F.E.M. and the condensed F.E.M. of the payload shall comply with the above general numbering 

requirement. 

4.6.1.5.2 Interface with the platform model 

This paragraph refers to the interface between the payload and the platform F.E.M. 

All co-ordinates are expressed in the satellite co-ordinate system defined §4.6.1.2.1. 

I/F Node name 

Co-ordinates (m) 

Grid 

Number X Y Z 

P1 50001 1.070 0.430 -0.430 

P2 50002 1.070 0.430 0.430 

P3 50003 1.070 -0.430 0.430 

P4 50004 1.070 -0.430 -0.430 

Table 4.6-4: Co-ordinates of the payload-platform I/F nodes 

4.6.2 THERMAL MODELS 

The need of thermal mathematical model is mission dependent. 

In a standard approach, only respective ICD is required (payload thermal ICD containing the geometrical model for 

satellite analyses and platform thermal geometrical model for payload analyses). 

If necessary, the exchange of electronic model will be discussed case by case. 

4.6.3 CAD MODELS 

All Computer Aided Design data exchanges between the Satellite Contractor and the Payload Supplier shall be 

based on CATIA V4.20 software .


4.7 SAFETY REQUIREMENTS 

PL - 4.5.7 -1 

deleted 

PL - 4.7 -1 

The Payload Supplier shall provide a safety analysis: 

• describing the hazardous items 

• identifying all hazardous events and associated causes 

• identifying all hazard controls and safety verification methods. 

• This analysis shall cover all phases from Payload Supplier delivery up to the launch site activities 

included. 

• This analysis shall include the GSE and operations. 

Nota : hazardous items can be pressurised items, pyrotechnic devices, ionizing and non ionizing radiation 

including lasers, batteries, lifting points, ignition sources… 

PL - 4.5.7 -2 

deleted 

PL - 4.7 -2 

The Payload shall be compliant with the Launch pad safety regulations in accordance with the contractual 

launch sites (depending on mission: launcher and launch site choice). 

PL - 4.7 -3 

Warnings and precautions relative to personnel and unit safety and hazards shall be specified in the payload 

handling, assembly, and test instructions. 

END OF CHAPTER 



PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: 1 

External Diffusion Sheet 

DIFFUSION CNES/PROTEUS 

Noms Sigles BPi Diffusion 

Action Information 

Ph. GILLEN (20 ex.) X 

External Companies 

Company Name Nb of Copies 



PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: 2 

Internal Diffusion Sheet 

Chef de Projet 5PF L. Frécon X P.A. Sureté/Fiabilité F.Cosson 

Chef de Projet Calipso M.Jourdan X P.A. Logiciel D.Lagelle 

Secrétariat M.Moreno X P.A. Sécurité D.Storto 

Process Manager P. Nicolas P.A. Composant Ph.Gasnier 

Controle Projet C.Bourget-Réné P.A. DHU J.F.Hernandez 

PA Manager G. Ferrier X P.A. Matériaux Procéd. E.Bordeux 

PA Adjoint V.Grossetete P.A. STR J.M.Cognet 

Configuration J.C. Daguzan P.A.Radiation O.Mion 

Contrat J. Pianca-Ripert X P.A. GPS / Roues C.Foggia 

Achat P.Borie VCF Ph.Cam 

Achat B.Durand BDS C.Lecrivain 

Resp. Tech. Proteus F.Douillet X AIT Ph.Chipon X 

Resp. Tech. 5PF T.Huiban X AIT système G.Obadia 

Ingénerie Sytème PF J.Camous X AIT avionique P.Ricca 

Resp. Tech. Calipso F.Paoli X AIT mécanique Patrice Moulin 

I/F Payload Calipso Y.Baillion X 

Architecte Cde & Ctrl S.Pouget X G.S R.Laget 

Cde & Ctrl DHU W.Medrecki OBSW LV L.Guibellini 

Gestion Bord Ph.Fourtier 

Architecte Electrique J.P. Canard X BANCS S.Vinay 

Architecte Elect.support E.Liebgott BANCS G.Nicolas 

Alimentation V.Michoud BANCS C.Bourgeois 

Harnais M.Preiti BANCS F.Maingam 

Alim. Batterie H.De Tricaud 

Alim.BEU. J.J.Digoin I.R.P. Avionique J.M.Bartolo 

PCE /Diode Box L.Gerreboo 

SEPTA (sadm) L.Canas 

EMC A.Luc 

Simulat.Energie J.F.Plantier 

Resp.Ch.Fonct. SCAO M.Sghedoni Direction Obs. et Sciences J. Chenet X 

Architecte SCAO F.Raissiguier X Directeur des prog de sat Sc 

et Obs de la Terre 

P. Mauté X 

Ingenerie SCAO J.L.Beaupellet Affaires futures Sc P. Kamoun X 

Ingenerie SCAO D.Brethé Ingénierie des Sat. Et PF – B.Lafouasse/ H. Sainct X 

Aff. Futures 

(7 ex) 

Ingenerie SCAO O.Rouat Ingénierie des Instruments – 

Aff. Futures 

JB Ghibaudo X 

GPS / STR J.L.Ribet 

GYR / CSS H.Dauphin 

MTB/MAG C.Lawrence 

RW T.Demas 

Propulsion T.Weulersse 

IRP Module D.Franqueville 

Structure J.Mourey 

Architecte AMT C.Duplay X Ingénieur Système P. Terrenoire X 

Aménag. CAO B.Cyvoct Ingénieur Système Y. Durand X 

AMT Analyses R.Knockaert Ingénieur Système JM. Nakache X 

Thermique M.Valentini X 

Etudes Modules (Ids/Icd) P.Laurenti 

DOCUMENTATION 

(original) 

X 



PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: i 

PROTEUS USER'S MANUAL 

This document contains the technical information which is necessary to: 

1. assess the compatibility of a payload with a PROTEUS platform 

2. assess the compatibility of a mission control centre with a PROTEUS satellite control centre 

3. prepare all technical and operational documentation related to a mission based on a PROTEUS 

system 

Missions out of PROTEUS standard flight envelope could be possible depending on launcher, orbit and 

payload parameters combination. In such a case, a more detailed analysis shall be done. 

This document will be revised periodically, comments and suggestions on all aspects of this manual will 

be encouraged and appreciated. 

Any questions concerning commercial aspects or interpretation of this manual should be directed to: 

Jocelyne PIANCA-RIPERT 

ALCATEL SPACE 

Proteus Sales Manager 

ALCATEL SPACE 

26, avenue Jean-François CHAMPOLLION 

BP 1187 

31037 TOULOUSE Cedex 1 

FRANCE 

Tel: 33 (0)5.34.35.46.68 

Fax: 33 (0)5.34.35.51.90 

Jocelyne.pianca-ripert@space.alcatel.fr 

Inquiries concerning technical clarifications of this manual should be directed to: 

Christian TARRIEU Francis Douillet 

CNES Proteus Program manager ALCATEL SPACE 

Proteus technical coordination and futur 

studies 

CNES 

BPi 2532 

18 Av. E.Belin 

31401 TOULOUSE CEDEX 4 

FRANCE 

ALCATEL SPACE 

100 Boulevard du Midi 

BP 99 

06322 CANNES LA BOCCA CEDEX 

FRANCE 

Tel: (33) 05.61.27.30.50 Tel: (33) 04.92.92.61.41 

Fax: (33) 05.61.28.13.21 Fax: (33) 04.92.92.79.50 

Christian.Tarrieu@cnes.fr Francis.Douillet@space.alcatel.fr 



PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: ii 

FOREWORD 

The PROTEUS development program is carried out under the partnership of the French Space Agency 

Centre National d’Etudes Spatiales (CNES) and ALCATEL SPACE. 

The equipment for PROTEUS system are provided by the industrial companies from countries such as 

Belgium, Canada, France, Germany, Italy, Spain, Sweden, USA. 

ALCATEL SPACE/CNES - ALL RIGHT RESERVED 

All information contained in the PROTEUS User’s Manual are proprietary to ALCATEL SPACE and 

CNES and shall be treated as such by the recipient party. It is supplied in confidence and shall not be used 

for any purpose other than the evaluation of PROTEUS capacities, and shall not, in whole or in part be 

reproduced, communicated or copied in any form or by any means (electronically, mechanically, 

photocopying, recording, or otherwise) to any person without priAll information contained in the 

PROTEUS User’s Manual are proprietary to ALCATEL SPACE and CNES and shall be treated as such by 

the recipient party. It is supplied in confidence and shall not be used for any purpose other than the 

evaluation of PROTEUS capacities, and shall not, in whole or in part be reproduced, communicated or 

copied in any form or by any means (electronically, mechanically, photocopying, recording, or otherwise) 

to any person without prior written permission from ALCATEL SPACE and CNES. 

Such right to use proprietary information shall not be deemed to imply any transfer or licence on 

intellectual property rights to such proprietary information, including patent, trademark, copyright, ideas, 

know how, methods or industrial design. 

ALCATEL SPACE and CNES could not held be responsible for the possible PROTEUS evolutions and 

evolutions of the launch vehicles compatible with PROTEUS based satellites. 

ALCATEL SPACE is ready to update the information concerning the launch vehicles capacities upon 

request based on Launch Service Agencies new data. 



PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: iii 

1. CNES 

CNES AND ALCATEL SPACE OVERVIEW 

Copyright CNES / C. Bardou / D. Ducros, 1997 

The Centre National d’Etudes Spatiales (CNES) is the French space agency. This public institution of industrial 

and commercial nature was founded in December 1961 in order to develop French space activities. 

The CNES role is to propose the directions that French space policy should take and, along with its partners 

(in industry research, and defence), to implement the programmes selected. 

CNES leads French space policy in two complementary ways: 

by playing a major role in European Space Agency (ESA) programmes, 

and by a dynamic national programme guaranteeing industrial competitiveness on a world level. 



PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: iv 

CNES is strongly linked with many different partners: space users, French industry, scientific laboratories, 

defence corps It also cooperates with foreign space agencies to fulfil ambitious programmes mainly within the 

realm of science. 

The major CNES programmes are those which involve the major strategic and economic challenges: 

access to space, with the Ariane programme and the creation of a launch base in French Guyana. 

Ariane is an ESA programme, whose launch services are marketed by Arianespace, 

space applications such as Earth observation (Spot, Topex-Poseidon, Jason, Polder/Adeos, Scarab, 

Vegetation,and so on...) and telecommunications (Telecom 2, Stentor, GNSS ...), 

science programmes in conjunction with research corps and led on the basis either European or 

international co-operation (Rosetta, intervention in Cassini-Huygens, Iso, Soho, Integral, Mars Sample 

Return,...), 

activities related to microgravity research and mankind in space (Alice2, Fertile, Castor,...) and the 

preparation of experiments designed for the International Space Station (Pharao), 

activities linked to Defence programmes (Helios, radar satellites ...). 

In order to fulfil its function, the CNES has various centres: Head Office in Paris, the launch vehicle directorate 

in Evry (responsible for the Ariane programme), the technical and operational centre in Toulouse (responsible 

for preparing and developing space projects for satellites and planetary vehicles as well as for running 

operational facilities and test infrastructures), the Guyana Space Centre and a balloon launch centre located 

in Aire-sur-l’Adour. CNES employs a total of 2500 staff spread throughout these five sites. 

The CNES budget stands at 12309 MF (almost 2052 M$, or 1876 MEuros),broken down into a State subsidy of 9265 

MF (almost 1544M$, or 1412 MEuros) and the Establishment’s own resources of 3044 MF (almost 507 M$, or 464 

MEuro). Over the past fifteen years or so, CNES has founded commercial subsidiaries to sell products and services 

arising from space technology. The 19 companies thus created directly employ a total staff of over 1000. 

CNES - Centre spatial de Toulouse 

18 Avenue Edouard Belin 

31401 Toulouse Cedex 4 - France 



PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: v 

2. ALCATEL SPACE : THE CANNES FACILITY 

The French industrial company ALCATEL SPACE was created by the merger of Aerospatiale Satellites 

(Cannes), Alcatel Espace (Nanterre, Toulouse), Thomson-CSF-spatial ground systems section (Buc, Toulouse, 

Kourou, Evry), Sextant avionique, spatial section (Valence), Cegelec (Telemetries activities at Kourou and in 

metropolitan France). ALCATEL SPACE holds leadership positions in all areas of satellite applications : 

Telecommunications (Arabsat, Eutelsat, Turcksat, Nahuel, Thaicom, Sinosat, Astra), navigation, observation 

(Helios, Vegetation), meteorology (MSG), science (ISO, Huygens) with a range of platforms (Spacebus, 

PROTEUS, Meteosat), payloads, instruments, products (microwaves, electronics, optics, radar, mechanisms, 

structures, thermal control...), ground segments, ground products and logistic support. ALCATEL SPACE owns 

partners throughout the world: strategic partners with Space Systems/Loral in the USA and partnerships or 

agreements with leading industrialists world wide (Europe : Matra Marconi Space, Alenia, Dasa - Canada : 

Spar - Japan : Toshiba, Mitsubishi, Sharp - Russia : NPO-PM - USA : Lockheed Martin, Hughes). ALCATEL 

SPACE 1998 turnover (forecast) amounts to 10 billion FF. (1.5 billionEuros) ALCATEL SPACE counts a 6000 

workforce. 

Cannes Centre has a staff of 1,300, more than 60% of whom being highly skilled engineers and 

professionals, making it thus stand out as the French Riviera's leading industrial employer. Over three 

decades of space activity, they have contributed to making this Operations Centre the Number1 European 

manufacturer, taking an active part in delivering over a hundred satellites to date. 

The Cannes Centre's research departments, laboratories and integration clean rooms are staffed by top notch 

specialists in numerous fields, from mechanics to electronics, from telecommunications to optics, from power 

supply to cryogenics. 

This multidisciplinary approach enables ALCATEL SPACE to provide complex systems, from inception to 

completion, in full compliance with Customer specifications, jointly with many French and foreign partners. 

The Cannes Centre's integrated environmental test facilities offer complete testing of satellites. It houses, in 

particular, Europe's largest space-oriented integration clean room for the complete assembly of up to six 



PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: vi 

satellites at a time, including a Compact Radiofrequency Simulator for measuring the radioelectric 

performance of antennas and satellites. And finally, there are powerful techniques for the integration and 

testing of spatial optical systems to produce increasingly sophisticated instruments. 

ALCATEL SPACE is rising to the challenges of the third millennium. In space and on the Earth we provide the 

tailor made solutions requested by our Customers. 

ALCATEL SPACE 

Cannes Center 

BP 99 - 06156 Cannes-la-Bocca Cedex - France 

Tel : (+33) 04 92 92 70 00 - Fax : (+33) 04 92 92 33 10 

http://www.alcatel.com 

350 SATELLITES 



PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: vii 

USER'S MANUAL CONFIGURATION CONTROL SHEET 

ED. REV. DATES 

MODIFIED 

PAGES 

CHANGES APPROVAL 

1 0 30/10/95 First Edition F. DOUILLET 

2 0 24/06/98 all 

3 draft1 30/10/98 all 

3 draft 2 30/11/98 all 

3 15/02/99 all 

Updating of the previous issue and insertion of new 

chapters from the draft CNES LDP.MU.L0.SC.300.CNES 

- Updating of the previous issue for chapters 1 and 2. 

- Deletion of the chapters 3 which presents detailed 

platform design. Instead of this chapter, Payload 

characteristics and satellite interfaces are presented. 

- Insertion of chapters 4, 5, 6 which deal with the payload 

interfaces, environment, verification tests. These chapters 

become the baseline to specified the studied mission 

requirements. 

- Chapter 7 is the ex chapter 5 

- Updating of chapter 8 which corresponds to the previous 

chapter6 

- Updating of the previous issue mainly chapters1,2 and 3. 

- Insertion of a new chapter between the ex chapter 7 and 

the previous chapter 8. 

- Updating of the previous issue 

- Insertion of a new chapter between the previous chapter 

7 and the previous chapter 8 



J. BLOUVAC 

B. LAZARD 

C. GRIVEL 

J. BLOUVAC 

C. GRIVEL 

J. BLOUVAC 

C. GRIVEL 

P. TERRENOIRE

PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: viii 


MODIFIED 

PAGES 

4 16/12/99 all 


TOTAL UPDATING 

- Updating with respect to : 

• ASPI-1999-ILF-142 

• ASPI-1999-ILF-150 

• ASPI-1999-ILF-160 

• ASPI-1999-ILF-162 

• ASPI-1999-ILF-170 

• ASPI-2000-ILF-006 

- Restructuration of the chapters 3 and 4 : 

• Chapter 3 becomes the description of the interface 

requirements 

• Chapter 4 becomes the description of the payload 

design requirements 

- Numbering of the requirements 

- Modification of the document in order to transform it in 

an applicable document for the payload 

- Addition of the IDS format and help in appendix 

- Addition of figures : 

• electrical interface brackets 

• STA interface 

- Chapter 1 

Updating with respect to the modifications of the other 

chapters. 

Introduction of section 1.7 (applicable & reference 

documents) and of section 1.8 (acronyms) 

Addition of the Star Tracker reference frames 

- Chapter 2 

Updating of the lower inclination for the allowable 

orbits 

Clarification for the visibility duration (addition of 

figures with 10° of elevation instead of 5°) 

Updating of the section 2.3.3 

... 



Y. BAILLION

PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: ix 


MODIFIED 

PAGES 

4 16/12/99 all 


Chapter 3 

It becomes the Payload interface requirements. 

Addition of some figures (electrical interface brackets, 

STA interfaces) 

Modification of the mass properties requirements 

Modification of the stiffness requirements 

Modification of the in flight allowable volume 

Clarification of the mechanical interfaces 

Modification of the maximum generated disturbances 

requirements 

Description of the active thermal control algorithm 

Addition of the « Power Supply requirements » section 

Total updating of the command / control sections 

Description of the active thermal control algorithm 

Addition of the « Power Supply requirements » section 

Total updating of the command / control sections 

Description of the pins allocation 

Addition of some information about the STA 

Addition of a « Ground Support Equipment 

requirements » section 

- Chapter 4 

It become the Payload design requirements. 

Updating of mechanical design requirements 

Addition of the « mathematical models interfaces 

requirements » section 

- Chapter 5 

Updating of the mechanical environment 

Updating of the sine environment 

Addition of a random environment 

Clarification of the shock requirement 

Modification of the thermal environment 

Modification of the magnetic field requirement 

Addition of the ground, storage and transportation 

environment 

- Chapter 6 

New organisation of the chapter. 

Introduction of new requirements about the payload 

instrumentation for satellite tests 

- Chapter 7 

No modifications 

- Chapter 8 


- Chapter 9 


- Chapter 10 

Identification of the delivering items responsibility 

Addition of an appendix containing the IDS files and an 

help for filling these IDS 



Y. BAILLION

PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: x 

ED. REV. DATES MODIFIED 

PAGES 


4 1 24/03/00 P ii, foreword modification, Rid PUM.4.0.BL.003 

P ix, change notice evolution 

P xii, TBC list evolution 

P xiii, TBD list evolution 

- Chapter 1 

P 1.13, figure 1.3-6, Rid PUM.4.0.YB.021 

P 1.13, S-band data rate, Rid PUM.4.0.BL.005 


P 1.14, Table 1.3-1, Rid PUM.4.0.BL.005 

P 1.16, Table 1.3-2, Rid PUM.4.0.BL.005, Rid 

PUM.4.0.YB.027 

P 1.16, section 1.3.5, Rid PUM.4.0.CG.003 

P 1.18, section 1.3.5.4, Rid PUM.4.0.CG.003 

P 1.19, section 1.3.5.4, Rid PUM.4.0.CG.003 

P 1.25, SY-1.4-8, Rid PUM.4.0.YB.002 

P 1.26, addition of figure, Rid PUM.4.0.YB.002 

P 1.28, SY-1.5-1, Rid PUM.4.0.YB.001 

P 1.29, Figure 1.5-1, Rid PUM.4.0.YB.001 

P 1.33, RD10 deleted, Rid PUM.4.0.CG.001 

P 1.34, addition of acronyms, Rid PUM.4.0.CG.003, 

Rid PUM.4.0.YB.012 

- Chapter 2 

P 2.18, last paragraph, Rid PUM.4.0.FD.002 

P 2.19, Table 2.4-2, Rid PUM.4.0.CG.005 

P 2.28, first sentence, Rid PUM.4.0.YB.026 

P 2.32, fig 2.5-14, Rid PUM.4.0.YB.040 

P 2.33, fig 2.5-16, Rid PUM.4.0.YB.003 

P 2.39, typing error, Rid PUM.4.0.YB.015 

… 



PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: xi 


PAGES 


Chapter 3 

P 3.2, , Rid PUM.4.0.FD.002 

P 3.3, PL-3.1.1-1, Rid PUM.4.0.YB.008 

P 3.4, PL-3.1.1-2 & 3, Rid PUM.4.0.YB.008 


P 3.4, Fig 3.1-1, Rid PUM.4.0.YB.045 

P 3.5, Fig 3.1-2 and 3.1-3, Rid PUM.4.0.FD.003 

P 3.6, PL-3.1.1-6 & 7 & 9, Rid PUM.4.0.YB.008 

P 3.6, PL-3.1.1-9, Rid PUM.4.0.YB.038 


P 3.2 to 3.9, equipped payload, Rid PUM.4.0.YB.008 


P 3.17, wording, Rid PUM.4.0.FD.004 

P 3.17, clarification, Rid PUM.4.0.YB.009 


P 3.21, text below PL-3.1.4-11, Rid PUM.4.0.BL.011 


P 3.24, table 3.2-1, Rid PUM.4.0.YB.018 


P 3.26, above section 3.2.2.1, Rid PUM.4.0.FD.005 



P 3.30, information addition, Rid PUM.4.0.YB.005 

P 3.30, information addition, Rid PUM.4.0.FD.007 

P 3.34, section 3.3.3.2, Rid PUM.4.0.BL.012 

P 3.34, section 3.3.3.3, Rid PUM.4.0.YB.039 


P 3.37, Figure 3.4-1, Rid PUM.4.0.FD.022 

P 3.38, level 3 definition, Rid PUM.4.0.YB.023 

P 3.38, 1553 time line addition, Rid PUM.4.0.YB.025 

P 3.39, PL-3.4.3-5 deleted, Rid PUM.4.0.YB.006 

P 3.39, PL-3.4.3-6, Rid PUM.4.0.YB.023 

P 3.40, PL-3.4.3-15 deleted, Rid PUM.4.0.YB.006 




PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: xii 


PAGES 


… 

- Chapter 3 (Continued) 


P 3.43, TBC deleted, Rid PUM.4.0.YB.014 


P 3.52, section 3.4.5.3.1.2, Rid PUM.4.0.YB.026 



P 3.57, addition of PL–3.4.7-4, Rid PUM.4.0.FD.011 



P 3.63, PL-3.5.3-1 and Fig 3.5-7, Rid PUM.4.0.FD.020 

P 3.63, monitoring frequency, Rid PUM.4.0.YB.044 


P 3.66, PL-3.5.3-10 deleted, Rid PUM.4.0.FD.011 

P 3.66, PL-3.5.3-11, Rid PUM.4.0.FD.011 

P 3.68, Figure 3.5-9, Rid PUM.4.0.FD.013 



P 3.87, wording, Rid PUM.4.0.YB.016 



P 3.98, section 3.5.8.2, Rid PUM.4.0.YB.007 

P 3.102, figure 3.6-3, Rid PUM.4.0.BL.007 

P 3.103, section 3.6.2.2.5, Rid PUM.4.0.BL.002 

P 3.104, PL-3.6.2-7, Rid PUM.4.0.YB.028 

P 3.105, PL-3.6.2-8, Rid PUM.4.0.BL.008 

P 3.106, section 3.6.5, Rid PUM.4.0.YB.030 

- Chapter 4 

P 4.8, PL-4.2.1-2, Rid PUM.4.0.YB.001 

P 4.9, PL-4.2.2-5, Rid PUM.4.0.FD.014 

P 4.10, Fig 4.2-2, Rid PUM.4.0.YB.043 

P 4.15, PL-4.3.2-2, Rid PUM.4.0.Jde.002 

P 4.20, Typing error, Rid PUM.4.0.YB.004 



P 4.34, section 4.6.1.2.3, Rid PUM.4.0.YB.029 

P 4.37, first paragraph, Rid PUM.4.0.YB.029 

- 



PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: xiii 


PAGES 


- Chapter 5 

P 5.2, Table 5.1-1, Rid PUM.4.0.YB.034 

P 5.3, Table 5.1-2, Rid PUM.4.0.YB.035 

P 5.5, fig 5.1-1, Rid PUM.4.0.JDe.001 


P 5.10, table 5.6-1, Rid PUM.4.0.YB.042 

P 5.12, section 5.8, Rid PUM.4.0.YB.017 

P 5.16, random vibrations, Rid PUM.4.0.YB.041 

P 5.18, random vibrations, Rid PUM.4.0.YB.041 

- Chapter 6 

P 6.2, last paragraph, Rid PUM.4.0.BL.016 


P 6.10, below PL-6.1.6-2, Rid PUM.4.0.JDe.003 

P 6.17, addition of PL-6.1.8-28, Rid PUM.4.0.YB.011 





- Chapter 7 

No modification 

- Chapter 8 

P 8.4, section 8.3.2, Rid PUM.4.0.BL.013 

P 8.6, Rid PUM.4.0.BL.014 

P 8.7, Rid PUM.4.0.BL.014 

P 8.8, Rid PUM.4.0.BL.014 

P 8.9, Rid PUM.4.0.BL.014 

P 8.12 to P8.22, Rid PUM.4.0.BL.014 

P 8.26, section 8.3.2, Rid PUM.4.0.BL.018 

P 8.27 to P 8.33, Rid PUM.4.0.BL.014 

P 8.36, Rid PUM.4.0.BL.014 

- Chapter 9 

P 9.3, S-band data rates, Rid PUM.4.0.BL.005 

- Chapter 10 

P 10.4, typing errors, Rid PUM.4.0.FD.016 



- Appendix A 

P 4, Mass definition, Rid PUM.4.0.YB.001 

P 13, Typing error, Rid PUM.4.0.FD.005 



PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: xiv 


4 2 23/06/00 

MODIFIED 

PAGES CHANGES APPROVAL 

1.33 

2.8 

2.13 

2.14 

2.15 

2.16 

2.17 

3.4 

3.5 

3.21 

3.22 

3.27 

3.30 

3.31 

3.32 

3.33 

3.33 

3.40 

3.40 

3.41 

3.45 

3.45 

3.46 

3.52 

3.57 

3.63 

3.64 

3.71 

3.72 

3.74 

3.75 

3.76 

3.77 

3.77 

3.78 

3.79 

3.79 

3.80 

3.80 

3.81 

3.83 

3.84 

3.85 

3.93 

3.93 

3.95 

3.95 

3.95 

3.96 

Addition of RD10 and RD11 

Addition of critical points for vertical configuration 

Yaw steering figures introduction 



Making-up 

Making-up 

PL-3.1.1-3 modification 

Typing error 


Section 3.1.5.1 clarification 

Figure 3.2-2 addition 

TBC suppression 


Figure 3.3-1 & 3.3-2 updates 

Figure 3.3-3 update 

PL-3.3.2-1 clarification 


Status word description clarification 

Command types clarification 

Timing requirement clarification 

1553 errors description 

PL-3.4.4-12 addition 


Description of bit 12 of pps signal 

Control of relays clarification 

Typing error 

HLC input voltage update 

LLC DHU output & User input updates 

CS16 DHU output & User input updates 

Clarification 

Analog telemetry measurement chain accuracy 

Th. acquisition DHU output & User input updates 

Thermistors acquisition measurement chain accuracy 

Digital relay electrical interfaces updates 

Section numbering modification 

Digital bi-level electrical interfaces updates 


AS16 electrical interfaces updates 

AS16 electrical interfaces updates 

Clock signal electrical interfaces updates 






Figure 3.5-25 deleted 

Figure 3.5-26 recalled Figure 3.5-25 




PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: xv 


MODIFIED 

4 2 23/06/00 3.96 

3.98 

3.99 

3.100 

3.101 

3.103 

3.104 

3.105 

3.105 

3.106 

3.107 

3.110 

4.9 

4.9 

4.11 

5.3 

5.8 

5.12 

6.36 

6.37 

6.38 

7.16 

Chap 10 

Append. C 

PAGES CHANGES APPROVAL 

Vandenberg RF environment TBC suppression 

Maximum magnetic moment TBC suppression 

Reference to Appendix C 

Figure 3.6-2 modification 

Figure 3.6-2b introduction 


Figure 3.6-4 introduction 


Section 3.6.2.3.2 addition 

Section 3.6.4 modification 

Molecular cleanliness provided 

Molecular cleanliness provided 




Sine environment confirmation 

TBC suppression (PL-5.4-2) 

Typing error 

TBD suppression (thermal vacuum tests) 



Typing error 

Total update 

Addition of appendix C 



PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: xvi 


PAGES CHANGES 



APPROVAL 

5 0 23/11/01 All pages Logo of Alcatel Space Industries updated I.Bénilan 

i Missions out of PROTEUS standard flight envelope: RID 

N°PUM.4.2.IB.015 

xiii Change notice evolution 

xviii TBC list evolution 

xix TBD list evolution 

xxi Table of Contents updated with "First page" and "Ch.0" 

Chapter 1 I.Bénilan 

1.12 Figure 1.3-5: RID N°PUM.4.2.IB.007 

1.16 Table 1.3-2: RID N°PUM42.IB.039. 

The range of possible payload masses is also modified 

as a consequence of the new STA mass (RID 

N°PUM42.YB.003 and impact on PL-3.1.1-1) 

1.26 Figure with the STA Reference Frame modified as in 

RID N°PUM.4.2.YB.003 (Figure 3.6-1), numbered and 

named 

1.27 SY-1.4-9: RID N°PUM.4.2.IB.048 

1.29 Figure 1.5-1: RID N°PUM.4.2.IB.047 

1.33 RD5: RID N°PUM.4.2.IB.057 

1.37 Addition of the acronym "w/o" 


2.19 Table 2.4-1: "Platform Inertias and Cog position in 

satellite co_ordinate system" replaced by "Platform 

Inertias in CoG Satellite Reference Frame and CoG 

position in Satellite Reference Frame" 

2.24 Figure 2.5-3: RID N°PUM42.IB.052 

2.39-40 §2.5.7.1 and Figure 2.5-19: RID N°PUM42.IB.039

PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: xvii 


PAGES CHANGES 



APPROVAL 


3.2 § between Ch.3 and 3.1: point added to close the 

sentence. 

3.2 § 3.1: RID N°PUM42.IB.001 

3.3 PL-3.1.1-1: mass allocated to the Payload modified as 

a consequence of the new STA mass (RID 

N°PUM42.YB.003) 

3.3 Table 3.1-1: RID N°PUM42.YB.003 and .IB.039 

3.4 PL-3.1.1-2: RID N°PUM42.IB.058 

3.4-6 §3.1.1.2 and in particular PL-3.1.1-3 + Figures 3.1-1, 

-2, -3: RID N°PUM42.IB.004 

3.7 PL-3.1.1-6, -7, -8, -9, -10: RID N°PUM42.YB.008 






3.16 Figure 3.1-11: RID N°PUM42.YB.009+ .IB.008 

3.17 Figure 3.1-12: RID N°PUM42.YB.009 

3.19 After PL-3.1.4-7: RID N°PUM42.YB.013 



3.23 Creation of §3.1.4.3.2.2: RID N°PUM42.YB.003 

3.24 PL-3.1.5-4 and Table 3.1-3: RID N°PUM42.IB.005 

3.26 PL-3.2.1-1: RID N°PUM42.YB.009 

3.26 §3.2.1: "on" is added after "shown" in the sentence 

"The Solar Array dimensions are shown Figure 3.1-7." 

3.26 Before Figure 3.2-1: RID N°PUM42.IB.020 (MLI 

thickness) 

3.27 Creation of Figure 3.2-3: RID N°PUM42.YB.007 


3.29 § between 3.2.2 and 3.2.2.1: RID N°PUM42.YB.005 

3.29 PL-3.2.2-3: RID N°PUM42.YB.001 

3.30 PL-3.2.2-5: RID N°PUM42.YB.011 

3.31 § 3.2.2.2: RID N°PUM42.IB.054 

3.30 Before and in PL-3.2.3-1: RID N°PUM42.YB.011 

3.34 Typing error: " suppressed. 

3.35-36 Figure 3.3-1, -2 and -3: RID N°PUM42.IB.026 

3.37 §3.3.3.3 and PL-3.3.3-2: RID N°PUM42.IB.040 

3.38 After PL-3.4.1-1: RID N°PUM42.IB.029 

3.39 Table before PL-3.4.1-2 numbered and named. 

3.39 PL-3.4.1-2: RID N°PUM42.IB.037 

3.41 Figure in §3.4.3 numbered, named and updated as a 

consequence of RID N°PUM42.IB.029. Typing errors 

also corrected. 

3.43 Figure in §3.4.3.4 numbered and named. 

3.44 Table of §3.4.3.4: RID N°PUM42.FD.01

PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: xviii 


PAGES CHANGES 

3.44 Table of §3.4.3.4: empty line before added. Table is 

numbered and named. 

Bullets of the subadresses are modified. 

3.45 2nd table of §3.4.3.4 numbered and named. 

3.48 PL-3.4.4-2: "be send" is replaced by "be sent" (typing 

error) 

3.51 §3.4.4.3.2 and PL-3.4.4-11: RID N°PUM42.FD.01 

3.52 PL-3.4.5-1: RID N°PUM42.IB.029 

3.52 PL-3.4.5-2: RID N°PUM42.IB.029 

3.53 Typing error after PL-3.4.5-4: "OBS" replaced by 

"OBSW". 

3.53 §3.4.5.3: RID N°PUM42.IB.029 

3.58 §3.4.5.4: RID N°PUM42.IB.029 

3.64-66 Figures 3.5-4, -5 and -6 and Table 3.5-1, -2 and -3 

issued from the updated Appendix B. 

3.67 PL-3.5.3-1: RID N°PUM42.YB.001 

3.68 Before PL-3.5.3-2: RID N°PUM42.IB.014 

3.68 PL-3.5.3-4: RID N°PUM42.YB.001 

3.69 § 3.5.3.3.1 1st line: RID N°PUM42.YB.001 

3.71 PL-3.5.3-19: RID N°PUM42.IB.013 

3.74 Table 3.5-4: RID N°PUM42.YB.012 

3.76-77 Modification of Table 3.5-7 and creation of Figure 

3.5-11a: RID N°PUM42.YB.012 

3.77 Before Figure 3.5-12: RID N°PUM42.YB.012 

3.78 Creation of Table 3.5-8a: RID N°PUM42.YB.012 

3.81 After Table 3.5-11: RID N°PUM42.YB.012 

3.82 Table 3.5-15: RID N°PUM42.YB.011 

3.83 Typing error in Table 3.5-16: "reciever" replaced by 

"receiver" 

3.83 Table 3.5-17 and under the table after "Protocol": RID 

N°PUM42.YB.012 

3.83 Table after "Protocol" numbered and named. 

3.85 Creation of Table 3.5-20a: RID N°PUM42.YB.012 

3.88 § 3.5.6.3.1 In orbit pulse duration: RID 


3.89 Creation of PL-3.5.6-17: RID N°PUM42.YB.014 

3.89 §3.5.6.5: RID N°PUM42.IB.029 

3.92 Fig 3.5-21: RID N°PUM42.YB.001 

3.94 § 3.5.7.1.2 c): RID N°PUM42.YB.001 

3.96 PL-3.5.7-11: RID N°PUM42.IB.030 

3.96 Figure after PL-3.5.7-11 numbered and named. 



APPROVAL

PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: xix 


PAGES CHANGES 

3.101 Figures after PL-3.5.8-1 and -2 numbered and 

named. 

3.102 PL-3.5.9-1 inherits from PL-5.7.2-1: RID 

N°PUM42.IB.060 

3.102 Creation of PL-3.5.9-3 from PL-5.7.2-2: RID 


3.102 PL-3.5.9-2 inherits from PL-5.7.1-1 : RID 

N°PUM42.IB.060 and .IB.012 

3.102 Creation of PL-3.5.9-4 from PL-5.7.1-2: RID 


3.103 §3.6.1 and Figure 3.6-1: RID N°PUM.4.2.YB.003 

3.104 Point added at the end of PL-3.6.2-1 

3.104 Figure 3.6-2: RID N°PUM.4.2.YB.003 

3.105 Figure 3.6-2b and §3.6.2.2.1: RID N°PUM.4.2.YB.003 

3.106 Figure 3.6-3: RID N°PUM42.IB.008 and .YB.003 

3.106 §3.6.2.2.3 and §3.6.2.2.4: RID N°PUM.4.2.YB.003 

3.106-107 §3.6.2.2.5 and PL-3.6.2-6: RID N°PUM.4.2.YB.003 

3.108 Point added at the end of PL-3.6.2-7 


3.110 PL-3.6.2-10 + Table 3.6-3: RID N°PUM42.IB.005 

3.110-112 §3.6.3: RID N° PUM42.YB.004 

3.112 §3.6.4 and PL-3.6.4-1: RID N° PUM42.YB.003 



APPROVAL

PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: xx 


PAGES CHANGES 



APPROVAL 


4.9 PL-4.2.2-2 and creation of Figure 4.2-0: RID N° 

PUM42.YB.006 

4.12 PL-4.2.3-1 and -3: RID N° PUM42.IB.044 

4.14 Table after PL-4.2.5-2 numbered and named 

4.33 Table in §4.5.1.1 numbered and named 

4.40 Tables in §4.6.1.3.1.1 numbered and named 

4.42-43 Formulae re-written with "Equation Microsoft 3.0" but 

unchanged w.r.t. PUM42 

4.45 Table in §4.6.1.5.1 numbered and named 

4.45 Table in §4.6.1.5.2 numbered and named 



5.9 Introduction to Table 5.6-1 and Figure 5.6-3 modified 

as a consequence of RID N°PUM42.IB.016 

5.10 Table 5.6-1: RID N°PUM42.IB.016 


5.12 PL-5.7.1-1, -2, PL-5.7.2-1 and -2 are deleted and 

moved to §3.5.9: RID N°PUM42.IB.060 

5.16-18 Tables and figures numbered and named 


6.1 "requirement" replaced by "requirements" 

6.11 Table 6.1-2 renumbered in 6.1-3 (there is already a 

Table 6.1-2 at page 6.7). PL-6.1.6-5 modified in 

consequence. 

6.12 PL-6.1.6-7: RID N°PUM42.IB.006 

6.12 Creation of §6.1.4 and PL-6.1.6-8: RID 


6.16 Table after PL-6.1.8-9 numbered and named 

6.20 PL-6.1.8-15: RID N°PUM42.YB.001 

6.21 Tables after PL-6.1.8-19 numbered and named 

6.36 Typing error in PL-6.2.3-2: "shallbe" replaced by "shall 

be" 

6.36 Table of §6.2.2 numbered and named

PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: xxi 


PAGES CHANGES 



APPROVAL 

Chapter 7 J-M.Touraille 

New issue of Ch.7: RID N° PUM.4.2.IB.051 

Chapter 8 J-M.Touraille 

8.1 "PROTEUS Generic Ground System" replaced by 

"PROTEUS Generic Ground Segment" 

8.2 In the title and in §8.1: "PROTEUS Generic Ground 

System" replaced by "PROTEUS Generic Ground 

Segment" 

New issue of Ch.8: RID N° PUM.4.2.IB.051 


9.2 "PROTEUS User Manual" replaced by "PROTEUS User's 

Manual" 

9.3 Table numbered and named 




10.8 §10.1.4: RID N°PUM42.IB.049 

10.13 §10.2.4.2: RID N°PUM42.IB.049 

10.16 §10.3.4: RID N°PUM42.IB.049 

10.17 PL-10.3.7-1: RID N°PUM42.IB.010 

10.18 PL-10.3.7-3: RID N°PUM42.IB.056 

10.18 PL-10.3.8-3: RID N°PUM42.IB.055 

Appendix A 

RID N°PUM42.YB.002 

Appendix B 

The whole appendix: RID N°PUM42.YB.002 

Sheets "Title" and "Drawings": RID N°PUM42.IB.002 

Sheets "Connectors": RID N°PUM42.IB.009 

Appendix C 

The whole appendix: RID N°PUM42.YB.003 

Y.Baillion 

Y.Baillion 

Y.Baillion

PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: xxii 


PAGES CHANGES 

All pages 

Alcatel Space instead of Alcatel Space Industries 

6 0 03/03/03 

i 

xviii 

xxv 

xxvi 

Front pages 

Responsibilities modified RID.PUM 5.0.FD.01 

diffusion list modified RID.PUM 5.0.FD.01 

Alcatel space technical contact modification RID.PUM 

5.0.FD.01 

Introduction 

CNES and Alcatel Space contacts changes RID.PUM 

5.0.FD.01 

Change notice evolution 

TBC list evolution 

TBD list evolution 

Chapter 1 

1.10 Figure 1.3-3 modified (Proteus evolution) RID.PUM 

5.0.FD.04 

1.12 Figure 1.3-5 modified (Battery Li Ion indicated), text 

below : battery Li Ion instead of battery Ni Cd : 

RID.PUM 5.0.FD.04 

1.14 Table 1.3-1 : 99.864 kbits/s instead of 24.562 kbits/s 

for low TM rate : RID.PUM 5.0.CG.09 

1.16 Table 1.3.-2 : updated performances : RID.PUM 

5.0.FD.04 

1.25 SY-1.4-6 & SY-1.4-7: typing error (Star instead of 

Start) : RID.PUM 5.0.FD.02 

1.26 Figure 1.4-3 modification due to STR modification 

RID.PUM 5.0.FD.02 + RID.PUM.5.0.CG.02 

1.29 SY-1.5.1 : launcher adapter definition added RID.PUM 

5.0.FD.05 

1.32 CAD and NASTRAN version update RID.PUM 

5.0.FD.02 

1.34 RD12 and RD13 introduction: debris analysis RID.PUM 

5.0.FD.10 

1.34 Section addition for reference of standards used in this 

document RID.PUM 5.0.FD.02 



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PAGES CHANGES 

Chapter 2 

2.2 Added sentence to precise the launch vehicles 

characteristics are given for information only RID. 

PUM.5.0.FD.09 

2.3 Table 2.1-1 updated for 4 launchers + mention “for 

information only” added. RID.PUM.5.0.FD.09 

2.4 Figure 2.2-1 modified (flight domain 600 km) 

RID.PUM.5.0.FD.06 

2.5 2.2.2.1 : environment explanation about 600 km limit 

added RID.PUM.5.0.FD.06 

2.2.2.2, 2.2.2.3, 2.2.2.4 : precisions about flight 

domain limitations added RID.PUM.5.0.FD.06 

2.45 2.5.8 debris analysis introduction RID.PUM.5.0.FD.10 

Chapter 3 

3.4 to 3.6 Figures 3.1-1, 3.1-2 and 3.1-3 modified (MCI Proteus 

evolutions) RID.PUM.5.0.CG.04 

3.11 PL-3.1.3-2 modified : PF height = 1070 mm instead 

of 1047 mm RID.PUM.5.0.FD.02 

3.12 Figure 3.1-7 modified (Proteus evolutions: column 

height) RID.PUM.5.0.CG.04 

3.13 Figure 3.1-8 update RID.PUM.5.0.CG.04 








3.23 PL-3.1.4-11 biases clarification RID.PUM.5.0.FD.02 

3.26 PL-3.1.5-4 shock level generated by the PL at PL/PF I/F 

modified RID.PUM.5.0.CG.05 

3.26 PL-3.2.1-1 +XS MLI blancket efficiency 


3.30 PL-3.2.2-5 : thermistors type clarification 




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PAGES CHANGES 

3.33 PL-3.2.3-1: thermistors type clarification + nb of 

Fenwal and Rosemount updated RID.PUM.5.0.FD.02 

3.34 PL-3.3.1-1 clarification RID.PUM.5.0.FD.02 

3.34 3.3.1: after PL-3.3.1-1 specification , power values are 

guaranteed at minimum RID.PUM.5.0.CG.05 

3.36 PL-3.3.2-1 et PL3.3.2-2 : 900 W peak power + TBD 

mission dependent RID.PUM.5.0.CG.05 

3.37 Figure 3.3-4 correction RID.PUM.5.0.FD.02 

3.40 3.4.2 initializing phases precision 

3.4.2 transitions performed “after operational 

coordination” added 

3.4.2 automatic transition description transferred to 

3.4.6.1 so paragraph corresponding deleted in this 

introduction 


3.41-3.42 3.4.3 1553 introduction completed +figure 3.4-4 

modified 

RID.PUM.5.0.FD.02, RID.PUM.5.0.FD.07 & 

RID.PDIS.5.0.FP.05 

3.43 PL-3.4.3-20 + figure 3.4-6 introduction 

RID.PUM.5.0.FD.02, RID.PUM.5.0.FD.07 & 


3.43 Paragraph 3.4.3.2 

PL-3.4.3.7, PL-3.4.3.8, PL-3.4.3.10, PL-3.4.3.11 

deleted 

PL-3.4.3-9 modified 


3.44 PL-3.4.3-13, “message types” instead of “types” 


3.45 After PL-3.4.3-17, 2. status word : bit 15 is not used 


3.46 Table 3.4-2: 6 th column title modified 


3.46 Table 3.4-2: for transmitter shutdown & override 

transmitter shutdown modifications 


3.49 PL-3.4.4-1 + text deleted 


3.51 PL-3.4.4-5 is modified + text added 

RID.PUM.5.0.FD.02- RID.PDIS.5.0.FP.07-RID.PDIS.5. 

0.FP.08 

3.51 PL-3.4.4-6 is completed 


3.52 PL-3.4.4-8 is modified 

RID.PUM.5.0.FD.02-RID.PDIS.5.0.FP.09 



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D. REV. DATES MODIFIED 

PAGES CHANGES 

3.52 Text deleted after PL-3.4.4-9 


3.52 For PL-3.4.4-9 text clarification 


3.53 PL-3.4.4-10 requirement clarification 


3.53 After PL-3.4.4-10 text clarification 


3.56 PL-3.4.5-5 completed and nota added 


3.57 PL-3.4.5-6 modified RID.PUM.5.0.FD.02 


and nota added RID.PUM.5.0.FD.07 

3.57 PL-3.4.5-8 nota added RID.PUM.5.0.FD.07 


3.58 PL-3.4.5-14 added RID.PUM.5.0.FD.07 

3.59 3.4.5.1.3.2: Introduction modified RID.PUM.5.0.FD.07 

3.59 3.4.5.1.3.3: Introduction modified RID.PUM.5.0.FD.07 

3.59 After PL-3.4.5-12: text modified RID.PUM.5.0.FD.07 

3.59 After PL-3.4.5-12: text suppressed RID.PUM.5.0.FD.07 

3.59 After PL-3.4.5-12: text added RID.PUM.5.0.FD.07 

3.63 3.4.6.1 introduction: SHM precision added (text 

suppressed in 3.4.2 introduced here) 


3.63 3.4.6.1 introduction : 2 lines may be ON in SHM 


3.64 PL-3.4.6-4 deleted RID.PUM.5.0.FD.02 

3.64 PL-3.4.6-5 introduction : Payload switch off in case of 

system monitoring by HW leading to SHM 



system monitoring by SW leading to SHM 



payload anomaly RID.PUM.5.0.FD.02 

3.66 After PL-3.4.7-1, text added : Pps available when GPS 

ON RID.PUM.5.0.FD.07 

3.66 PL-3.4.7-3: 825 ms instead of 875 ms 


3.70 –3.72 PL –3.5.2-1 description H01, H02, H03 modified : 

figures 3.5-4, 3.5-5 et 3.5-6+tables 3.5-1, 3.5-2, 

3.5-3. RID.PUM.5.0.CG.06 

3.75 PL-3.5.3-7: deleted. Information only 


3.76 PL-3.5.3-8: deleted. Information only 


3.80 PL-3.5.4-4: added. RID.PUM.5.0.FD.03 



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PAGES CHANGES 

3.80 PL-3.5.4-5: added. RID.PUM.5.0.FD.03 

3.81 PL-3.5.6.1: duration between consecutive HLC for 

information only. RID.PUM.5.0.FD.02 

3.81 PL-3.5.6.2: Table 3.5-4 modified: fault voltage 33 V 

RID.MUP.5.0.CG.03 

3.82 PL-3.5.6.3: Table 3.5-5 modified: input voltage 21.5 

V-fault voltage 33 V RID.MUP.5.0.CG.03 

3.83 PL-3.5.6.4: Table 3.5-6 modified: diff.output 

impedance RID.MUP.5.0.CG.03 

3.83 PL-3.5.6.4: Table 3.5-7 modified: single input to 

ground RID.MUP.5.0.CG.03 

3.85 PL-3.5.6.6: Table 3.5-8 modified: diff.output 

impedance RID.MUP.5.0.CG.03 

3.85 PL-3.5.6.6: Table 3.5-9 modified: single input to 

ground RID.MUP.5.0.CG.03 

3.88 Table 3.5-11 modified: fault voltage (tolerance) 


3.88 Table 3.5-12 modified: fault voltage (emission) 


3.89 Table 3.5-15 correction : thermistors type and 

resistance RID.MUP.5.0.FD.02 

3.91 Table 3.5-18 modified RID.MUP.5.0.CG.03 

3.91 Table 3.5-20 modified: diff.output impedance 


3.92 Table 3.5-21 modified: single input to ground 


3.109 PL-3.5.9-1 modified payload ON and OFF separated 

RID.MUP.5.0.FD.02 

3.109 PL-3.5.9-2 modified RID.MUP.5.0.CG.08 

3.109 PL-3.5.9-4 modified RID.MUP.5.0.CG.08 

3.109 Tables 3.5-25 et 3.5-26 : volume vhere B is maximum 

introduction RID.MUP.5.0.CG.08 

3.111 Figure 3.6-1 modified RID.MUP.5.0.CG.02 


3.113 Figure 3.6-2b modified RID.MUP.5.0.CG.02 


3.113 STR mass modified RID.MUP.5.0.CG.02 

3.114 3.6.2.2.3 centre of gravity information deleted, see 

appendix C RID.MUP.5.0.FD.02 

3.114 3.6.2.2.4 moments of inertia information deleted, see 


3.114 3.6.2.2.5 moments of inertia information deleted, see 


3.114 3.6.2.2.5 stiffness RID.MUP.5.0.CG.02 



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PAGES CHANGES 

3.115 PL-3.6.2-2 correction RID.PDIS.5.0.FP.031 

3.115 PL-3.6.2-3 correction RID.MUP.5.0.FD.02 




3.115 Text added :real azimuth angle clarification 



3.117 PL-3.6.2-11 corrected RID.MUP.5.0.FD.02 

3.117 PL-3.6.2-8 corrected RID.MUP.5.0.FD.02 

3.117 Text added after PL-3.6.2-9 RID.MUP.5.0.FD.02 

3.117 Table 3.6.2 modified RID.MUP.5.0.CG.02 

3.118 PL-3.6.3-1 modified and text suppressed 


3.120 PL-3.6.3-3 updated RID.MUP.5.0.FD.02 

3.120 After PL-3.6.3-3, text suppression RID.MUP.5.0.FD.02 

3.120 After PL-3.6.3-3 text addition on STR cable 

characteristics RID.MUP.5.0.FD.02 

3.120 Introduction of paragraph 3.6.4 : Zsta MLI blanket 

efficiency provided RID.MUP.5.0.FD.02 

3.121 Introduction of paragraph 3.7 : TBD added (additional 

requirements on GSE) RID.MUP.5.0.FD.02 

3.121 PL3.7.1-1 completed RID.MUP.5.0.FD.02 

3.121 PL3.7.1-1 modified + nota :PF handling points 


3.121 3.7.1.2 introduction clarification 


3.123 PL3.7.1-10 completed RID.MUP.5.0.FD.02 

3.123 PL3.7.1-10 corrected RID.MUP.5.0.CG.04 

3.123 Section 3.7.1.2.4 added RID.MUP.5.0.FD.02 

3.124 PL-3.7.2-2 completed RID.MUP.5.0.FD.02 

3.129 PL-3.7.2-18 completed RID.MUP.5.0.FD.02 



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PAGES CHANGES 

Chapter 4 

4.10 PL-4.2.2-4 corrected RID.MUP.5.0.CG.04 

4.10 PL-4.2.2-5 modified-RID.MUP.5.0.CG.04 

4.10 PL-4.2.2-6 deleted-RID.MUP.5.0.CG.04 

4.11 Figure 4.2.1 corrected RID.MUP.5.0.CG.04 

4.12 Figure 4.2.2 corrected RID.MUP.5.0.CG.04 

4.13 PL-4.2.2-7 deleted RID.MUP.5.0.CG.04 

4.13 Text and PL-4.2.2-8 introduction RID.MUP.5.0.CG.04 

4.13 Introduction of paragraph 4.2.3 completed- 


4.13 PL-4.2.3-1 to PL4.2.3-3 requirements update- 


4.15 PL-4.2.5-2 clarification-RID.MUP.5.0.FD.02 

4.15 PL-4.2.5-7 pressurised item added- 


4.15 After PL-4.2.5-3 qualification loads clarification 


4.34 PL-4.4.5-3 deleted RID.MUP.5.0.FD.02 

4.37 Nastran version update RID.MUP.5.0.FD.02 

4.42 Table 4.6-1 typing error correction 


4.48 Paragraph 4.7 “safety requirements “introduced – 


Chapter 5 

5.4 Table 5.1-4 updated RID.MUP.5.0.CG.04 

5.8 PL-5.4-1 maximum pressure value modified- 


5.12 PL-5.7.2-1 typing error correction RID.MUP.5.0.IB.01 

5.18 Table 5.11-7 hoisting and handling clarification 


Chapter 9 

9.3 Low bit rate at 99.864 kbit/s RID.MUP.5.0.CG.09 

Appendix C 

Standard STA ids update-RID.MUP.5.0.CG.02 



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6 1 03/03/03 

PAGES CHANGES 

Chapter 0 

xxvii Table of contents modified (impact of 

RID.PUM.5.0.CG.10) 

Chapter 2 

2.37 Figure 2.5-18 changed for solar activity profile 

+introduction text modified RID.PUM.5.0.CG.01 

2.38 Table 2.5-1 corrected + text introduction modified 

RID.PUM.5.0.CG.01 

Chapter 3 

3.3 Table 3.1-1 updated RID.PUM.5.0.FD.04 

3.4 Center of gravity height = 0.73 m instead of 0.75 m 

RID.PUM.5.0.CG.04 

3.75 Introduction of 3.5.3.3 section modified 

RID.PUM.6.0.FP.29 

3.95 Pps characteristics update (just before PL-3.5.6-15) 


Chapter 4 

4.10 Handling attach fittings instead of handling attached 

fittings, RID.MUP.5.0.CG.04 

Chapter 5 

5.12 5.8 meteroid typing error correction 

RID.MUP.5.0.IB.01 

Appendix A 

Ids filling rules update RID.MUP.5.0.CG.10 

Appendix B 

Payload ids update-RID.MUP.5.0.CG.10 

Appendix D 

Platform Ids introduction RID.MUP.5.0.CG.10 



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APPROVAL 

6 2 09/07/04 Chapter 0 E. JAUFFRAUD 

Table of contents modified 

Chapter 1 E. JAUFFRAUD 

Configuration Control Sheets are displayed at the 

beginning of the chapter 






















Appendix B E. JAUFFRAUD 

Payload ids replaced by Payload Platform ids- 

RID.CIIS.4.1.JC.4_4 

Appendix D E. JAUFFRAUD 

Platform Ids ireplaced by STR user’s manual 

RID.CIIS.4.1.JC.1_13

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PAGES CHANGES 



APPROVAL 

6 3 22/11/04 Chapter 0 E. JAUFFRAUD 

One Configuration Control sheet added 



beginning of the chapter – List of TBCs/TBDs added 



beginning of the chapter – List of TBCs/TBDs added 



beginning of the chapter - List of TBCs/TBDs added 













Appendix A O. LERONDE 

In relation with Payload Platform IDS update 

Appendix B O. LERONDE 

Payload Platform IDS updated -RID.PUM.6.2.OL.02

PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: xxxii 

TBC / TBD list 

Tables are displayed at the beginning of each chapter. 



PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: xxxiii 


First page 

Ch.1 Scope 

Contacts - Foreword - CNES & ASPI Overview 

Configuration Control Sheet 

Ch.2 Mission envelope 

Ch.3 Payload interface requirements 

Ch.4 Payload general design requirements 

Ch.5 Payload environment requirements 

Ch.6 Payload verification and test requirements 

Ch.7 Generic PROTEUS control ground segment 

Ch.8 PROTEUS Generic Ground System (PGGS) - 

Mission Centre interfaces 

Ch.9 On board - ground interfaces 

Ch.10 Standard schedule, deliveries and documentation 



PRO.LB.0.NT.003.ASC issue 06 rev. 03 Page: xxxiv 

Appendix A IDS Filling Rules 

Appendix B Payload/Platform IDS 

Appendix C Standard STA IDS 

Appendix D STA User’s Manual 




Chapter 1 : Scope 


Here below are listed the changes between issue N-2 and issue N-1 


§1.3.3 Modified in useful TM data flow rate PUM.6.1.CG.06 

§1.5 [SY - 1.5 - 1 a] Modified in STR cable introduction PUM.6.1.JC.1_1 

§1.8.2 New in Introduced in J2 IISs PUM.6.1.EJ.32 










§1.8.2 Modified in modified in J2 IISs PUM.6.1.EJ.32 








Here below are listed the changes from the previous issue N-1 


§1.4 [PL - 1.4 -3 ] New in Antenna boresight transfer matrix to be 

provided 

PUM.6.2.EJ.02 

All right reserved ALCATEL SPACE/CNES 

Passing and copying of this document , use and communication of its content is not permitted without prior written authorization



CHANGE TRACEABILITY Chapter 1 1 

1. Scope 5 

1.1 PURPOSE OF PROTEUS USER’S MANUAL 5 

1.2 SERVICES 8 

1.2.1 PROTEUS STANDARD SERVICES 8 

1.2.2 EXTENDED SERVICES 8 

1.2.2.1 Specific adaptations 8 

1.2.2.2 Payload Instrument Module 8 

1.2.2.3 Payload 8 

1.2.2.4 Launch vehicle procurements 8 

1.2.2.5 Mission ground segment 8 

1.2.2.6 In orbit operations 9 

1.2.2.7 Full turnkey system 9 

1.2.3 INDUSTRIAL SHARING 9 

1.3 SYSTEM OVERVIEW 10 

1.3.1 SYSTEM ARCHITECTURE 10 

1.3.2 THE SATELLITE SYSTEM CHARACTERISTICS 11 

1.3.3 GENERAL PLATFORM DESCRIPTION 13 

1.3.4 PROTEUS MAIN CHARACTERISTICS 19 

1.3.5 PROTEUS BASED SATELLITE MODES 19 

1.3.5.1 Satellite OFF mode 21 

1.3.5.2 Satellite Start-up 21 

1.3.5.3 Satellite Test Mode 22 

1.3.5.4 Satellite Safe Hold Mode (SHM) 22 

1.3.5.5 Satellite Star Acquisition mode (STAM) 23 

1.3.5.6 Satellite Normal Mode 23 

1.3.5.7 Satellite OCM Modes (OCM2 and OCM4) 23 

1.3.6 GROUND CONTROL SEGMENT CHARACTERISTICS 23 

1.4 FRAMES AND SATELLITE AXIS DEFINITION 26 

1.5 DEFINITIONS 32 

1.6 UNITS, MODELS AND CONSTANTS 34 

1.6.1 UNITS 34 

1.6.2 MODELS 34 

1.6.3 CONSTANTS 35 

1.7 REFERENCE AND APPLICABLE DOCUMENTS 36 

1.7.1 REFERENCE DOCUMENTS 36 

1.7.2 APPLICABLE DOCUMENTS 37 

1.7.3 STANDARDS 37 

1.8 ACRONYMS 38 

1.8.1 REQUIREMENTS ACRONYMS 38 

1.8.2 OTHER ACRONYMS 38 





Figure 1.3-1: PROTEUS system architecture........................................................................................................... 11 

Figure 1.3-2 : Typical satellite based on PROTEUS ................................................................................................ 12 

Figure 1.3-3 : internal lay out of the PROTEUS platform ........................................................................................ 13 

Figure 1.3-4 : PROTEUS platform overview ........................................................................................................... 14 

Figure 1.3-5 : PROTEUS functional block diagram ................................................................................................ 15 

Figure 1.3-6 : Payload data path.......................................................................................................................... 16 

Figure 1.3-7 : Telemetry flow................................................................................................................................ 17 

Figure 1.3-8: Satellite modes................................................................................................................................ 21 

Figure 1.4-1 : Local orbital reference frame.......................................................................................................... 27 

Figure 1.4-2 : Satellite Reference Frame (For information, JASON 1 Satellite: PROTEUS platform+JASON 1 specific 

Payload) ....................................................................................................................................................... 28 

Figure 1.4-3: STA Reference Frame ...................................................................................................................... 30 

Figure 1.5-1 : Satellite architecture ....................................................................................................................... 33 


Table 1.3-1 : Main data flows characteristics ........................................................................................................ 18 

Table 1.3-2: PROTEUS main characteristics .......................................................................................................... 19 






LIST OF TBCs 

. 

N° § Sentence Planned 

Resolution 

§1.3.2 The PROTEUS platform has been designed to be compatible with various orbits 

(phased, sun synchronous, frozen and inertial orbits) with altitudes ranging from 

500 km to 1500 km, for an orbital plane inclination contained between 20 (TBC) 

and 145 deg. Flight domain is detailed in chapter 2.2.1. Mission design support is 

provided in paragraph 2.4. 

§1.3.5.2 The satellite is powered, the receivers are ON (hot redundancy, not commandable, 

automatically ON at satellite powering) and the Reconfiguration Module (RM) is 

waiting for separation strap disconnection. It is possible to send direct TCs (TCD) to 

command ON a Processor Module (PM) or modify the RM registers which define the 

on board configuration which shall be used after Umbilical Strap disconnection. If a 

PM is set ON, this PM will nominally detect the Umbilical presence and go to Test 

Mode. Nominally, the payload is OFF. It is not powered or, in case of special 

needs, in a reduced way depending on launch phase (30 W maximum for 2 of the 

16 power lines which can be maintained ON (TBC) and which are not managed 

but protected with a passive system (fuses) (see section 3.5.3)). Its thermal control is 

not ensured by the platform and the satellite attitude is imposed by the launch 

vehicle. 

§1.6.3 Orbit bulletin will be given in adapted parameters (a, ex, ey, i, Ω, α) in Veis 

reference frame (TBC) with 

§1.8.2 TBC To Be Confirmed 

LIST OF TBDs 

. 

N° § Sentence Planned 

Resolution 

§1.8.2 TBD To Be Determined 




1. SCOPE 

1.1 PURPOSE OF PROTEUS USER’S MANUAL 

Chapter 1 

P roteus overview 

Proteus services 

S atellite S ystem characteristics 

P latform description 

Ground control segment 

Chapter 7 

Ground segment functions, 

architecture, 

operations concepts, data 

exchanges 

Proteus? 

For a first approach, this 

chapter could be skipped 

Chapter 8 

Mission centre/Ground segment 

interfaces described in details 

P roteus Ground segment ? 

Chapter 9 

On board/Ground 

interfaces 

On board/Ground 

interfaces? 

Chapter 10 

S chedule, deliveries, 

documentation for 

typical Proteus mission 

The User 

S chedule, deliveries, 

documentation? 

C hapter 2.1,2.2,2.3,2.4 

orbit types,pointings,satellite orientations 

launch vehicles possible with Proteus 

Proteus 

missions panel? 

mission 

parameters choice? 

Chapter 2.5 

notions for mission 

analysis 

Payload 

compatible? Chapter 3 

P ayload Interfaces 

P ayload design 

rules? 

R equirements 

Payload 

Chapter 4 

environment ? 

P ayload design and construction 

requirements 

Payload tests? 

Chapter 5 

Payload environment requirements 

Chapter 6 

P ayload verification and 

test requirements 




PROTEUS is a generic name for a multimission platform designed for low Earth orbits: « Plate-forme 

Reconfigurable pour l’Observation, les Télécommunications Et les Usages Scientifiques ». This platform and the 

associated ground control segment have been developed together by ALCATEL SPACE and the CNES . 

This document is intended to be a reference manual which presents the general capabilities offered by the 

PROTEUS system. Indeed, its purpose is to allow a User looking for an efficient way to access space in low Earth 

orbits, to assess different mission profiles and solutions to achieve his objectives, to design User payload 

compatible with PROTEUS bus and launch vehicles, to build and verify his payload within the constraints imposed 

by the satellite bus and launch vehicles. 

In order to do so, Chapter 1 gives an overview of the PROTEUS system and its main characteristics, so that the 

User can easily evaluate the efficiency of this kind of concept for the baselined mission. 

Chapter 2 is entirely devoted to PROTEUS capabilities and indicates the main mission options, grouping the 

potential launch vehicles, achievable orbits, possible pointing modes, and different orbit types. The User can then 

determine the required orbit kind, main orbital parameters, pointing mode, and launch vehicle which are 

compatible with the mission objectives. 

Chapter 3 describes the interfaces requirements for a payload based on PROTEUS platform. For mechanical 

thermal, electrical and command & control domains, the requirements at payload level are listed. The User, 

responsible for the payload (considered as one element) must read this chapter to check the payload compatibility 

with the standard PROTEUS platform. 

This chapter deals with other interesting points for the authority in charge of the payload : 

• the satellite operational features implying some constraints for the payload, 

• the star trackers assembly accommodation as star trackers are laid out on the payload, 

• the interfaces between the Ground Support Equipment (GSE) and the payload for the satellite integration 

and alignment phase. 

Chapter 4 defines mechanical, thermal, electrical and command & control requirements for payload design and 

construction. 

Chapter 5 presents payload requirements due to the flight environment imposed by the chosen launch vehicle, 

the mission environment parameters (mission objective, orbit kind, mission date and duration). The listed 

requirements are estimated taking into account the envelope of the launch vehicles compatible with PROTEUS; 

that means the flight and qualification levels for the payload, (and the satellite) could be reduced as soon as the 

considered launch vehicle envelope is restrained. In this chapter, payload requirements for ground operations, 

storage, transportation and handling phases are detailed too. 

Chapter 6 lists payload design verification tests before payload delivery and briefly presents tests and verification 

at satellite level. 

Chapters 3, 4, 5 and 6 are a suitable baseline to tailor the payload requirements to the satellite bus ones. The 

tailoring of these chapters to the studied mission is an efficient tool to gather the payload requirements, to write 

the Payload Design Interface Specification and to identify very early the points needing a specific analysis. 

Chapter 7 presents the generic ground segment for PROTEUS based satellites. This chapter describes the 

functions, the operations concepts and operational organisation, the architecture, the performances for the 

ground control segment. 

Chapter 8 gives in detail the information necessary to understand and handle data exchanged between Mission 

Centre (MC) and PROTEUS Generic Ground System (PGGS). For a first approach with PROTEUS based mission, 

the User could skip this chapter. 

Chapter 9 deals with the on board-ground interfaces, all characteristics of the communication links between the 

ground control/command station(s) and the platform. 

Chapter 10 presents the standard schedule, deliveries and documentation for a typical PROTEUS mission. 




In any case, given the very wide range of possibilities offered by the PROTEUS system, ranging from the assembly 

and delivery of the platform to a full turn key system, the User is strongly encouraged to contact ALCATEL SPACE 

or CNES to help him analyse the mission and design the most appropriate solution, according to his needs. 




1.2 SERVICES 

1.2.1 PROTEUS STANDARD SERVICES 

The PROTEUS standard services ensured by ALCATEL SPACE and CNES consist in providing: 

• the satellite platform, 

• the satellite engineering, assembly, integration and test, 

• the generic ground control segment procurement including a ground station and a control centre, 

• the transportation, the satellite launch campaign activities and the first operations including orbital 

checkout, 

• the control centre operations. 

1.2.2 EXTENDED SERVICES 

1.2.2.1 Specific adaptations 

Any requirement at platform, payload interfaces, mission levels not described in this document can be studied 

case by case by ALCATEL SPACE and CNES; for instance PROTEUS based mission may present smaller payload 

or heavier one, need more power... 

1.2.2.2 Payload Instrument Module 

ALCATEL SPACE and CNES propose a Payload Instrument Module based on a standard design compared with 

PROTEUS platform and easily adaptable. It allows: 

• either to integrate a payload composed of several boxes, 

• or to be a module between the platform and the main payload instrument; its function consists in 

containing various electronic boxes, harness to connect the payload instrument to the platform, and/or an 

optional X band data communication subsystem. 

1.2.2.3 Payload 

The wide field of activities and the important experience of ALCATEL SPACE make possible the delivery of specific 

payload instruments, according to the Customer’s needs. 

Adopting this functional scheme also allows to optimise the system activities as ALCATEL SPACE is involved at 

satellite platform level. 

1.2.2.4 Launch vehicle procurements 

ALCATEL SPACE or CNES can provide the launcher. This task covers all the interface management activities with 

the launch vehicle provider from the choice of the launch vehicle (about 2.5 years before launch) to the launch 

campaign itself. This activity includes all the necessary safety considerations. 

1.2.2.5 Mission ground segment 

The mission centre is usually specific to each mission. It can be developed by ALCATEL SPACE or the CNES upon 

request. 




1.2.2.6 In orbit operations 

CNES or ALCATEL SPACE can be in charge of operating PROTEUS command-control ground segment for some 

missions. CNES is also responsible for Launch and Early Operation Phase. Once in orbit, the operations 

(including station keeping) can also be performed by the CNES. This is a good way to reduce operations cost 

because the teams can work simultaneously on different missions. 

1.2.2.7 Full turnkey system 

As it is done for some geostationary telecommunications missions or in other domains, ALCATEL SPACE is able to 

provide a full turn-key system if desired by the Customer. 

1.2.3 INDUSTRIAL SHARING 

The industrial sharing on a PROTEUS based mission depends on the Customer’s needs. It typically depends on 

the origin of the contract: commercial bids versus governmental space agencies procurements. In the case of a 

commercial contract, ALCATEL SPACE can be the prime contractor, supplying the Customer with a complete 

operational system, including operations and ground control for instance. For scientific missions which result from 

an international co-operation like for Jason (USA/French co-operation), or for a CNES mission like for Corot 

another entity such as a national space agency is responsible or co-responsible for the mission and may want to 

take the responsibility for the command-control ground segment and operations. Then, ALCATEL SPACE is 

responsible for the platform, integration and test of the payload on the platform and for the satellite level 

engineering. Other tasks may depend on each mission. 




1.3 SYSTEM OVERVIEW 

PROTEUS offers a standard multimission platform for a very attractive cost and within a delivery time of 24 

months (from end of phase B (PDR) to launch for a standard mission). Technically, the platform architecture is 

generic. Adaptations are limited to relatively minor changes in a few electrical interfaces and software modules. 

The robustness and low cost properties of this recurring design concept have been demonstrated, and very 

different missions such as Jason 1 for radar altimetry, Picasso-Cena for earth environment, Corot for astronomy, 

and commercial for optical Earth and radar observation plan to use the PROTEUS system. 

1.3.1 SYSTEM ARCHITECTURE 

The architecture of a space system based on PROTEUS is shown on Figure 1.3-1. 

The central column gives the main components of a standard PROTEUS system: 

• the standard platform, 

• the standard control command ground system (SSGP) with the station (TTC-ET), the satellite control centre 

(CCC) and the ground network, 

• all documentation, hardware and software needed to produce and to test a satellite including payload 

instruments, 

• launch pad and first in-orbit operations. 

The left column describes the mission specific contribution to the system: 

• the payload instruments, 

• the mission control centre where the payload operations are commanded and the payload data is 

processed, 

• an optional TM station used to increase the visibility duration or the ground segment availability. 

The right column presents the launch vehicle chosen by the mission system manager and not included in the 

standard PROTEUS service, and one CNES 2 GHz station used to help the first acquisition after launch 

separation. 




Figure 1.3-1: PROTEUS system architecture 

1.3.2 THE SATELLITE SYSTEM CHARACTERISTICS 

The PROTEUS platform has been designed to be compatible with various orbits (phased, sun synchronous, frozen 

and inertial orbits) with altitudes ranging from 500 km to 1500 km, for an orbital plane inclination contained 

between 20 (TBC) and 145 deg. Flight domain is detailed in chapter 2.2.1. Mission design support is provided in 

paragraph 2.4. 

The platform with its folded solar arrays is compatible with small launch vehicle fairing internal diameters from 

1.9 m. 

The platform provides a wide range of payload pointing capabilities (Earth and anti-Earth pointing, inertial 

pointing); typical pointing performance is 0.05 deg (3σ). 

Satellite based on PROTEUS belong to the 500 kg class with a payload mass between 100 kg and 275 kg, 

consuming up to 300 W power. Typical satellite based on this platform is shown on Figure 1.3-2. 




Jason mission 475 kg/400W 

altitude = 1336 km/ inclination = 66 deg/Earth pointing 

Figure 1.3-2 : Typical satellite based on PROTEUS 




1.3.3 GENERAL PLATFORM DESCRIPTION 

Figure 1.3-3 and Figure 1.3-4 show the general lay out of a PROTEUS platform. 

.. 

Figure 1.3-3 : internal lay out of the PROTEUS platform 




Figure 1.3-4 : PROTEUS platform overview 

The platform structure has a 1 m sided cubic shape. All the equipment units are accommodated on four lateral 

panels; hydrazine mono-propellant units with a 40 litre tank and four 1 N thrusters are laid out on and under the 

lower plate. The interface with the launch vehicle is through an adapter (specific to each launch vehicle) bolted to 

the bottom of the structure. The mechanical interface with the payload is provided through four points at the 

corners of the upper panel. The platform features a structure with frame permitting panel removal and easy 

integration. When payload topology allows for it, the platform structure concept is reused for the payload module 

structure. 

The PROTEUS functional block diagram is shown on Figure 1.3-5. The functional redundancies are fully ensured 

at satellite level; as far as the hardware is concerned, the equipment units are either one-to-one, or n out-of m 

redundant (for example : 2 gyros out of 3, 3 reaction wheels out of 4...) 




9s3p 78A.h 

Li Ion Battery 

Figure 1.3-5 : PROTEUS functional block diagram 

The thermal control subsystem is dimensioned to withstand the maximum thermal loads defined by PROTEUS 

candidate mission. The concept relies on passive radiators and active regulation with heaters, monitored by the 

central computer. Mission adaptation is limited to MLI windows dimensioning and thermal control parameters 

adjustments. To ensure the safety and health of the satellite payload, PROTEUS provides thermal control and 

heater power to the payload in all satellite modes. 

Electrical power is generated by two symmetric wing arrays attached near to the satellite centre of mass with two 

single-axis stepping motordrives. Each wing is comprised of four 1.5*0.8 m panels covered with classical silicon 

cells. The power is distributed through a single non-regulated primary electrical bus (23/36V with an average 28V 

voltage), using a Li Ion battery (9s 3 p technology) developed by SAFT. 

The electrical, on-board command and data handling architecture is centralised on one single computer, the 

Data Handling Unit (DHU). Functionally, one half of the satellite is under the control of one processor within the 

DHU, and the other half of the satellite is under the control of the other processor. 

The primary functions devoted to the Data Handling Unit are: 

• Satellite modes management consisting of automatic mode transitions and routines. 

• Failure detection, isolation and recovery (FDIR), consisting of monitoring of satellite health and switching to 

safe hold mode if necessary. 

• On-board observability, consisting of generation, maintenance, and downlink of housekeeping telemetry 

data. 

• Satellite commandability, consisting of telecommands sent by ground either to hardware or software. 

The DHU performs most of its tasks through the central MA 31750 processor which runs the satellite software. It is 

responsible for the power distribution to all satellite units. 




It also supports the management of the communication links to each satellite unit either through discrete point to 

point lines or via MIL-STD-1553 B bus. The processor generates a clock reference, manages satellite data 

storage, and ensures telemetry frame decoding. A maximum of 1000 time-executable commands may be 

uplinked and stored in any given pass, although additional immediate commands can be sent during satellite 

ground visibility. 

The DHU command buffer can hold a maximum of 20 kwords (16-bit words) and is the constraining element in 

the uplink commanding capability. Payload commands are relayed to the payload at a maximum rate of 8 Hz. 

The DHU manages the payload throughout the commands, offers standard thermal control, and standardized 

electrical interfaces (23/36V power supply, 1553 bus, specific point to point lines). A payload specific software 

application can be implemented in the DHU to control complex payload. 

The Data Handling Unit has an internal mass memory organised in two main areas : 

• a housekeeping area (HKTM-R) to record payload and platform housekeeping data out of visibility periods. 

• a data payload area of 2 Gbits, split in two size programmable areas (PLTM1 / PLTM2) to store and 

transmit independently payload data to ground during visibility periods. 

The data is transmitted from payload to mass memory through a1553 link with a maximum rate of 100 kbits/s or 

through a specific high speed line with a data rate up to 10 Mbits/s (optional). 

A 722.116 kbit/s S band QPSK downlink (without encoding -reedsolomon nor Viterbi- nor frame packetting) is 

available for telemetry. A ground control capacity is provided by a 4 kbit/s S band up-link. The CCSDS packet 

standard protocol is used for TM encoding and TC decoding. 

Figure 1.3-6 shows the general payload data path and Figure1.3-7 shows more particularly the telemetry flow 

from the payload to the ground system. 

.. 

Figure 1.3-6 : Payload data path 




Figure 1.3-7 : Telemetry flow 

Table 1.3-1 gives the useful TM information flows (except for transport overhead) produced on board and 

transmitted to ground, the on-board rate characteristics used as basis to size the information transport and 

storage functions. The rate is the average over one day to allow easy calculation of the quantities of information 

to be stored and sent to the ground. 




Flow function Data rate characteristics 

(average over one day) 

HKTM-P knowledge of the 

satellite status and 

configuration 

HKTM-R detailed satellite 

surveillance 

Scientific 

PLTM 

information 

produced by the 

payload to be sent 

to mission center 

Observations Transmission 

to the ground 

N/A received during 

visibility if 

emitter is ON 

-300 bps when useful TM data 

rate is 85.966 kbit/s 

- 500 bps when data rate is 

722.116 kbit/s 

rate very variable 

depending on the 

mission phase and 

the ground 

programmation 

mission dependent the control centre 

does not know the 

packets contents 

-recorded on 

board 

-transmitted 

upon request 

from the ground 

-recorded on 

board 

-sent upon 

request from the 

ground 

Table 1.3-1 : Main data flows characteristics 

For more information, the chapter 9 deals with the on board / ground interfaces in details. 

Accurate attitude determination is based on two star trackers (nominal and redundant) measurements. Both star 

trackers are accommodated on the payload in a Star Tracker Assembly (STA) equipped with an autonomous 

thermal control. 

The normal in-orbit platform attitude control is based on a gyro-stellar concept. Three accurate 2-axis gyrometers 

are used for stability requirements and attitude propagation. Attitude acquisition is obtained using magnetic and 

solar measurements (two 3-axis magnetometers and eight coarse sun sensors). Platform attitude control can 

provide a rotation around the axis perpendicular to the solar array driving mechanisms (yaw steering), allowing a 

90 % recovery of sunlight in the case of a non sun synchronous orbit. 

Four small reactions wheels will generate torque for attitude command, and are de-saturated using magnetic 

torquers. 

A Global Positioning System (GPS) receiver will provide satellite position information for accurate orbit ephemeris 

determination and on board time delivery. 

The unavailability for a PROTEUS based satellite is estimated to 0.82 % with 0.25% due to reconfigurable failures 

(example : switch froman equipment unit to a redundant one), 0.50% due to the radiation effects, 0.07% due to 

the orbit correction manoeuvres. This last mission interruption case is calculated assuming manoeuvres of 15 

minutes/month (mission dependent). The Star Tracker occultation could imply a damaged pointing performance 

and so a mission unavailability, but nominally the star tracker lay out on the payload is optimised to avoid it. 




1.3.4 PROTEUS MAIN CHARACTERISTICS 

Table 1.3-2 summarises the main characteristics and related performances of the PROTEUS platform. 

Orbit any orbit altitude in 500-1500 km 

orbit inclination higher than 20 deg (TBC) 

Launch vehicles compatible with all launch vehicle 

with fairing diameter >1.9m 

Mass dry platform mass w/o STA = 262 kg 

28 kg hydrazine capacity 

Payload mass = 100 to 286 kg 

Reliability 0.892 over 3 years 

0.759 over 5 years 

Lifetime 3 to 5 years depending on the orbit 

Power bus maximum consumption = 300 W 

Payload consumption class = 200 W 

up to 300 W on some orbits 

Pointing 

Attitude restitution 

0.05 deg (3 σ) on each axis 

Data storage 2 Gbits for payload 

Down link 722.116 kbits/s 

Up link 4 kbits/s 

Unavailability 0.81 % 

Table 1.3-2: PROTEUS main characteristics 

1.3.5 PROTEUS BASED SATELLITE MODES 

Various satellite « modes » are used to define the behaviour of the satellite, together with the associated 

configuration of the equipment units, and their monitoring. 

The main in-flight modes are driven by the Attitude and Orbit Control System (AOCS) : 

• Start up mode 

• Safe Hold Mode (SHM) 

• Star Acquisition Mode (STAM) 

• Normal mode (Nom) 




• Orbit Correction Mode with 2 thrusters (OCM2) 

• Orbit Correction Mode with 4 thrusters.(OCM4) 

Two other modes are used on ground : 

• Off mode 

• Test mode. 

All these modes, shown on Figure 1.3-8 are described in more details hereafter. 

Notice : During transition phases (for instance manoeuvres for orbit correction depending on the mission and 

during Safe Hold Mode), the payload could be dazzled. 

An automatic transition to Safe Hold Mode (on redundant equipment) is automatically engaged after critical on 

board failure detection. If there is an instrument in failure, this instrument will be put in passive state and powered 

OFF; but it does not imply a satellite transition to Safe Hold Mode. 




Figure 1.3-8: Satellite modes 

Notice: In flight, the Safe Hold mode-Start Up mode transition duration is less than around 1.5 minutes. 

1.3.5.1 Satellite OFF mode 

The satellite is not powered (battery not connected), the satellite is not operable. This mode is used for storage or 

transportation. 

1.3.5.2 Satellite Start-up 

This mode is used during the launch and on ground. 




The satellite is powered, the receivers are ON (hot redundancy, not commandable, automatically ON at satellite 

powering) and the Reconfiguration Module (RM) is waiting for separation strap disconnection. It is possible to 

send direct TCs (TCD) to command ON a Processor Module (PM) or modify the RM registers which define the on 

board configuration which shall be used after Umbilical Strap disconnection. If a PM is set ON, this PM will 

nominally detect the Umbilical presence and go to Test Mode. Nominally, the payload is OFF. It is not powered 

or, in case of special needs, in a reduced way depending on launch phase (30 W maximum for 2 of the 16 

power lines which can be maintained ON (TBC) and which are not managed but protected with a passive system 

(fuses) (see section 3.5.3)). Its thermal control is not ensured by the platform and the satellite attitude is imposed 

by the launch vehicle. 

This mode is normally engaged in two ways : 

• either with the connection of the battery to the Power and Conditioning Equipment unit (PCE) ; in that case, 

it lasts from ground operations on the launch pad till separation from the launcher. 

• or on an alarm triggering ; in that case, the switch to start up mode is only a transient. 

Safe Hold Mode transition is automatically engaged after umbilical strap disconnection detection. 

The exit from this mode is performed in two ways: 

• either the automatic way when the satellite is separated from the launch vehicle, 

• or by a high priority ground command to enter the test mode; this transition is used on the launch pad, 

under ground control, when the satellite is connected via an umbilical cord, or during AIT. 

1.3.5.3 Satellite Test Mode 

This mode is used during AIT or on the launch pad for final verifications. On the launch pad, the allowed TCs are 

limited to the ones necessary for health checks. 

This mode is typically engaged with a high priority ground command (TCD) on the launch pad. 

The exit of this mode is performed by using a TCD on the launch pad: switch OFF Processor Module A or B. 

On the launch pad, it is possible from this mode to return to Start up Mode by a telecommand. 

1.3.5.4 Satellite Safe Hold Mode (SHM) 

This mode consists in 3 main phases : RDP (Rate Damping Phase), SPP (Sun Pointing Phase), BBQ (Barbecue). 

The aim of this mode consists in reaching autonomously a safe attitude with the -X satellite axis pointed to the Sun 

and with the mean roll angular rate equal to -0.25 deg/s. In this mode, thrusters and sophisticated equipment 

(for instance gyros) are not used. 

For Safe Hold mode transfer, in the first phase, the satellite changes its attitude in order to place its solar array in 

canonical position (40 min) and then is oriented until its -Xs axis points to the Sun. To achieve this attitude, the 

satellite can be briefly oriented such as the payload is dazzled by the Sun. 

Then, the satellite may stay as long as necessary in Safe Hold Mode. In this mode, PROTEUS provides the 

minimum amount of satellite management required to support vital functions for diagnosis or anomaly handling. 

These include ground to satellite communication, thermal control, battery management, failure management, 

reduced (30W) payload power (see section 3.5.3), and coarse sun pointing. Coarse sun sensors and 

magnetometers provide attitude measurement, and magnetic torquers generate torque. In addition, two of the 

four reaction wheels are used to provide gyroscopic stiffness. 

This mode is always engaged after the initialisation transition defined by the Start up Mode. This mode begins 

after the first initialisation, or on an alarm triggering, or on a software reset. 

The satellite leaves the Safe Hold Mode: 

• normally, when the OBSW identifies the TC which commands the Transition mode 




• if an alarm occurs, implying an automatic transition to Start up Mode. 

1.3.5.5 Satellite Star Acquisition mode (STAM) 

Star Acquisition Mode is used to reacquire fine attitude, position, and time information during the transition from 

Safe Hold Mode to Nominal Mode. These data are given by the Global Positioning System and Star Tracker. It 

starts when the satellite can be considered « safe ». 

In this mode, the payload operations are restricted to the ones strictly necessary to verify the instruments 

behaviour. Priority is given to housekeeping operations. Nominally payload is OFF. In case of special needs, 2 of 

the 16 power lines can be maintained ON which insure reduced (30W) payload power. Nominal payload 

thermal control is performed by the OBSW. 

The only way this mode is engaged is when receiving a ground command while on Safe Hold Mode. 

The mode exit is performed on ground request, with the TC mode change to Nominal Mode. 

1.3.5.6 Satellite Normal Mode 

The normal mode, in addition to the vital satellite management functions, provides generic or specific services as 

required by the payload, including power, commanding and status via 1553B bus, precise datation and fine 

pointing. 

The Normal mode is engaged in three different ways: 

• from the Star Acquisition Mode, upon ground request, 

• automatically when leaving the Orbital Correction Mode, 

• in a Normal Mode reset process, upon ground request; this method is used to change the attitude and 

orbit control system equipment configuration. 

Usually, the satellite stays in Normal Mode. An alarm or a ground request could involve a mode exit towards the 

Safe Hold Mode via the Start-up Mode. 

1.3.5.7 Satellite OCM Modes (OCM2 and OCM4) 

Orbit Manoeuvres are commanded in these modes. The performances and services are the same as in Nominal 

mode, but attitude pointing can be damaged and payload functioning can be restricted during large 

manoeuvres. The Satellite is under the OBSW control. The thrusters are used to control the orbit: 4 thrusters are 

needed when an important delta V is to be performed; only 2 thrusters are required when small manoeuvres are 

necessary. As the thrust direction is aligned with the +Xs satellite axis, the satellite shall be oriented such as the 

+Xs axis is aligned with the velocity vector during these manoeuvres. 

The periodicity and the choice of these modes depend on specific mission analysis and control. During these 

manoeuvres, payload units may be either operating or not, according to the mission. 

Those modes are always engaged upon ground request. The transition is always performed from the Nominal 

Mode. 

Exit from these modes is performed automatically : 

• when the orbit manoeuvre is over, to nominal mode 

• on alarm, to start up mode. 

1.3.6 GROUND CONTROL SEGMENT CHARACTERISTICS 

The ground segment called « PROTEUS Generic Ground Segment » (PGGS) consists of one or more 

Telecommand and Telemetry Earth Terminal (TTCET), a Command and Control Center (CCC) and a Data 




Communication Network (DCN). The mission centers (MC) are connected to the CCC and the TTCETs via the 

DCN. 

The TTCET consists in three subsystems : 

• the « radio frequency » subsystem sets up the on board/ground link on order from the CCC. It ensures the 

reception of the signals delivered by the satellite and the transmittal of signals to the satellite. Reception is 

made in circular polarization diversity mode and transmission according to a polarization selected by the 

CCC. 

• the « base band » subsystem receives the telecommands from the CCC and transfers them to the radio 

frequency subsystem. It transmits satellite telemetry to the CCC and to the MCs. Completely automated, it 

does not require the operators presence. 

• the « time frequency » subsystem distributes a time reference and a frequency reference to the radio 

frequency and base band subsystems. 

The CCC ensures the telemetry processing, satellite orbit and attitude control functions, the generation and 

transmission of platform telecommands, the reception and transmission of mission telecommands from the MC. 

The operating modes of the CCC depend on the needs of the User mission and range from partially automatic 

operation during working hours and on working days to all manual operations 24 hours-a-day. 

The consultation function for the archived data (TM parameters, raw TM packets, the satellite status, the logbook 

and the operational documentation) and distribution of results is performed by a WWW data server. The 

Customer uses either a standard workstation equipped with a navigator, or a specific station (DRPPC) equipped 

with a navigator and packages supplied by the CCC. 

The protocols used for PGGS data transfer are the following : 

• TCP-IP (Internal Protocol) for real time exchanges between the Command and Control Center (CCC) and 

the Telecommand and Telemetry Earth Terminal (TTCET) (housekeeping, telecommand, RC, RM) and for 

transmitting TTCET RMs to the Mission Center(s). 

• FTP (File Transfer Protocol) for files transfer (housekeeping, payload telemetry, pointing data, mission 

generation help data) 

• HTTP and e-mail for data exchanged between the CCC and the expert DRPPCs. 

For a given mission, the PROTEUS Generic Ground Segment (PGGS) is a part of the mission ground segment. It 

does not ensure all the mission functions but all those required for final orbit acquisition and control of the 

satellite. Its main functions are the following ones: 

• Satellite surveillance and technical control 

these functions consist in checking, thanks to housekeeping telemetry processing, that the status of the satellite is 

satisfactory for mission needs and for transmitting telecommands to maintain normal satellite operation. 

• Orbit and attitude control 

Satellite orbit determination is performed by the PGGS from the Global Positioning System (GPS) data received in 

housekeeping telemetry. If an orbit correction is required, the PGGS generates the control commands which are sent 

by telecommands and executed by the satellite. Attitude control is performed automatically on board from the GPS 

data and the sensor information. The PGGS periodically updates the attitude control on board model parameters. 

• Payload service 

This consists in transmitting the received payload telemetry to the Mission Center, checking the status of the payload 

thanks to housekeeping telemetry processing and in performing the programming operations at the frequency 

dependent on mission requirements and satellite storage capacity. 

• Satellite expert appraisal 




It consists in leading investigations in case of satellite misfunctioning or reports on its behaviour. These operations 

are led by the operators of the Control Center or by external experts. 




1.4 FRAMES AND SATELLITE AXIS DEFINITION 

Inertial Reference Frame J2000.0 

SY - 1.4 - 1 

This reference frame shall be used at system level 

J2000 means the date 01/01/2000 at 12h00 (barycentric dynamic time) 

The X axis is the mean equinox of the J2000 date, the intersection between the mean equator of the J2000 date 

and the mean ecliptic of the J2000 date. 

The Z axis is colinear to the poles axis with the South-North direction. The poles axis is perpendicular to the mean 

equator of the J2000 date. 

The Y axis completes the right-handed orthogonal reference frame. 

Earth Reference Frame WGS 84 

SY - 1.4 - 2 


WGS 84 means World Geodesic System. 

For PROTEUS ground/on board calculations requirements, the WGS 84 Earth Reference Frame is considered as the 

same as the IERS (International Earth Rotation Service) reference frame: IRTF90, IERS terrestrial reference system for 

the year 90. 




Local Orbital Reference Frame (G, X Lo, Y Lo, Z Lo) F Lo 

SY - 1.4 - 3 

This reference frame shall be used at system level. 

It is known as the conventional pitch, roll and yaw system. 

G is the satellite center of mass in operational conditions. 

Y Lo (Pitch axis) is perpendicular to the orbital plane and oriented in the opposite direction of the orbital kinetic 

momentum. 

Z Lo (Yaw axis) is parallel to the orbital plane and oriented toward the geocentric direction. 

X Lo (Roll axis) completes the right-handed orthogonal reference frame (this axis is parallel to the velocity vector and 

oriented in the same direction if the orbit is perfectly circular). 


Passing and copying of this document , use and communication of its content is not permitted without prior written authorization 

+Z Lo 

+X Lo 

Figure 1.4-1 : Local orbital reference frame 

+Y Lo


Satellite Reference Frame (P, Xs, Ys, Zs) Fs 

SY - 1.4 - 4 


It is used to define hardware location within the satellite. 

P is located at the center of the launch vehicle interface circle: at the bottom of the standard PROTEUS interface 

frame and the top of the specific launch vehicle adapter. 

+Zs is parallel to the launch vehicle interface plane and pointed towards H01 electrical bracket (towards Earth in 

normal flight configuration). This axis defines the normal to the « Earth panel ». 

+Xs is perpendicular to the launch vehicle interface plane and oriented from launch vehicle towards satellite. 

+Ys completes this right-handed orthogonal reference frame (this axis is parallel to the launch vehicle interface 

plane and parallel to the solar array rotation axis). 

This frame is materialized by a reference mirror cube located on the platform bottom frame. 

The frame axis is shown with Jason payload as example on Figure 1.4-2. 

Figure 1.4-2 : Satellite Reference Frame 

(For information, JASON 1 Satellite: PROTEUS platform+JASON 1 specific Payload) 




Satellite Center of Gravity Reference Frame (G, X G, Y G, Z G) F G 

SY - 1.4 - 5 


It is obtained by a simple translation of the Satellite Reference Frame (Fs) to the satellite center of gravity (G). 

It is the reference for the satellite mass, centering and moments of inertia configuration. 

Star Tracker n°2 Frame (O STR2, X STR2, Y STR2, Z STR2) F STR2 

SY - 1.4 - 6 

This reference frame shall be used at system level and is defined by 

OSTR2 is the geometric centre of the reference mirror cube of Star Tracker 2 

XSTR2 is parallel to the Star Tracker 2 CCD length and is pointed towards Star Tracker 1 

Z STR2 is parallel to the Star Tracker 2 optical axis toward space 

Y STR2 completes this right-handed orthogonal reference frame ( it is parallel to the CCD width) 

Star Tracker n°1 Frame (O STR1, X STR1, Y STR1, Z STR1) F STR1 

SY - 1.4 - 7 


OSTR1 is the geometric centre of the reference mirror cube of Star Tracker 1 

XSTR1 is parallel to the Star Tracker 1 CCD length and is pointed in the same direction as XSTR2 Z STR1 is parallel to the Star Tracker n°1 optical axis toward space 

Y STR1 completes this right-handed orthogonal reference frame ( it is parallel to the CCD width) 

The reference for satellite attitude determination is reference mirror cube of Star Tracker 1. 

Star Trackers Assembly Frame (OSTA, XSTA, YSTA, ZSTA) FSTA SY - 1.4 - 8 


OSTA is geometric centre of the interface points of the STA 

XSTA is parallel to the STA interface plane and is in the same direction of XSTR1 Z STA is perpendicular to the interface plane and is pointed toward the payload 

Y STA completes this right-handed orthogonal reference frame ( it is parallel to the STA interface plane) 

The reference frame is shown on Figure 1.4-3. 




Payload Reference Frame (O P, X P, Y P, Z P) F P 

PL - 1.4 - 1 

Figure 1.4-3: STA Reference Frame 

This reference frame shall be used at payload level. 

It is obtained by a translation of the Satellite Reference Frame (Fs) to the Payload Interface Plane. 

O P is the geometric centre of the four platform/payload interface points in this interface plane. 

This frame is materialized by a reference mirror cube located on the Payload Instrument Module, on or close to 

the Star Tracker bracket. 

Instrument Unit Reference Frames 

PL - 1.4 - 2 

This reference frame shall be defined for each Instrument Unit and shall be documented in the corresponding IDS. 

This frame shall have preferably Z axis perpendicular to the unit mounting plane. 

For Instrument Units with accurate pointing constraints, this Instrument Unit reference frame will be materialized 

by a reference mirror cube. 




PL - 1.4 -3 

The transfer matrix from the instrument antenna boresight (if any) to its reference frame shall be provided in its 

mechanical IDS. 

Satellite Attitude 

SY - 1.4 - 9 

The satellite attitude is defined by the orientation of the Satellite Center of Gravity Reference Frame (G, XG, YG, 

ZG) or the Satellite Reference Frame (P, Xs, Ys, Zs), versus the Local Orbital Reference Frame (G, XLo, YLo, Zlo). 

In case of nominal configuration (see section 2.3.1.1.1.2 a), the Satellite attitude is defined by the Euler Angles, 

defined in the following order: 

Ψ: Yaw angle = positive rotation around ZLo (from XLo toward YLo) 

θ: Pitch angle = positive rotation around the image of YLo after the Ψ rotation. 

φ: Roll angle = positive rotation around Xs (image of XLo after Ψ and θ rotations). 

For any other configuration, these definitions will be mission dependent (vertical flight, inertial pointing,...). 

An equivalent representation of the attitude is given by the quaternion representation [q0, q1, q2, q3] with the 

convention q0>0 (real part) . 




1.5 DEFINITIONS 

These definitions are used to distinguish standard PROTEUS elements and mission specific elements making a 

satellite. 

SY - 1.5 - 1 a 

These definitions shall be used at system level 

Unit or Instrument Unit: A unit is a single box defined by an equipment/assembly name, an identification part 

number and a serial number. 

Instrument or Payload Instrument: An instrument or a payload instrument is defined by one unit or a set of 

units with associated harness necessary to perform a type of measurement or a payload function. 

Payload Instrument Module (PIM): The payload instrument module supports the payload instruments, and 

provides a thermally controlled environment and the necessary harness to connect the Payload instrument to the 

platform. 

Payload: The payload is the assembly of the Payload Instrument Module and the Payload Instruments. 

Equipped Payload: Payload + STA + H02 & H03 connectors brackets + STR cables 

STA : The Star Trackers Assembly is a part of the platform but shall be mounted on the payload. The assembly 

payload + STA + connectors brackets + STR cables is the equipped payload. 

Platform or PROTEUS Platform: The PROTEUS platform provides all the necessary housekeeping functions to 

perform the mission: Payload support, electrical power, command, data handling and storage, attitude and orbit 

control,... 

Launcher adapter: mechanical ring bolted on PROTEUS standard platform at one end and specific to the 

launcher clamp band at the other end, equipped with thermal protections and actuator interface pads. 

Bus or Satellite Bus: The Bus is the assembly of the PROTEUS Platform, the Payload Instrument Module (mission 

specific) and the launcher adapter (launcher specific) which stays attached to the platform after launch vehicle 

/satellite separation. 

Satellite: The satellite is the assembly of the Bus and the Payload Instrument or (equivalent) the assembly of the 

Platform, the Payload and the Launcher adapter. 

To allow an easy adaptation to missions, the interfaces between the platform and the payload are standardised 

and clearly defined. 

This architecture is illustrated in Figure 1.5-1. 

For missions with a single instrument, the Platform Instrument Module can be suppressed if the instrument fits with 

these standard interfaces. 




Figure 1.5-1 : Satellite architecture 




1.6 UNITS, MODELS AND CONSTANTS 

1.6.1 UNITS 

For PROTEUS studies, the physical parameters are expressed in metric system. 

SY - 1.6 - 1 

These units shall be used at system level. 

• Length in millimetre (mm), Metre (m) or kilometre (km), 

• Mass in kilogram (kg), 

• Inertia in kg x m 2 , 

• Pressure in Bars (b) or millibars (mb) or in Pascal (Pa) for very low pressure, 

• Temperatures in degrees Kelvin (K) or Celsius ( °C), 

• Thermal inertia in J/K, 

• Angles in degree (deg), arc minute and arc second, or in microradian (µrad) for small angles, 

• Specific impulse in Seconds (s), 

• Power in Watt (W). 

1.6.2 MODELS 

A PROTEUS based satellite is designed with the following models : 

CAD model: 

CAD model is developed on CATIA version 4.21. 

Structural model: 

Structural analysis will be performed on NASTRAN version or 70. 

Thermal model: 

Thermal analysis is performed on CORATHERM release 99. 

Radiations: 

The models used for the calculation of 

trapped electrons and protons are AE8 and AP8 models, NASA environment models, 

heavy ions ; LET curve is established with the Cosmic Ray Effects on Micro Electronics (CREME) programme. The M 

« Weather index » of CREME is usually equal to 3 which corresponds to « Galactic Cosmics Rays + Adams 90 % 

worst case Solar activity ». 

Earth magnetic field model: 

Reference: IGRF95 International Geomagnetic Reference Field 1995. Order: 10. 

Model IGRF is used. 

Atmospheric model: 

a) BARLIER : A thermospheric model based on satellite drag data. A-N°185-1997. 




b) monoatomic oxygen : In ESABASE, the software which permits to calculate the monoatomic oxygen effect is called 

ATOMOX. 

1.6.3 CONSTANTS 

Orbitography: 

Orbit bulletin will be given in adapted parameters (a, ex, ey, i, Ω, α) in Veis reference frame (TBC) with 

a = semi major axis in km 

ex = e x COS (perigee argument) = X component of eccentricity vector 

ey = e x SIN (perigee argument) = Y component of eccentricity vector 

i = inclination in deg 

Ω = inertial ascending node longitude in deg 

α = perigee argument + true anomaly = position on orbit (pso) in deg 

Rt 

Earth potential field: 

The model GEM10 will be used: 

= Earth ellipsoid equatorial radius = 6378.140 km 

G = Universal gravitational constant = 6672 x 10-14 kg-1 .m3 /s2 µ = Earth gravitational constant = 398600.64 km 3 /s 2 

a = Earth ellipsoid flatness = 0.003352836 

J2 = second zonal harmonic = -1.0826268 x 10-3 T = Earth sidereal period = 86163.9796 sec 

q1 = Earth sidereal rotation rate 

Other main constant: 

= 0.00417808 deg/sec 

c = light speed = 299792.458 km/s 

Ps = Solar pressure coefficient = 4.56 10-6 Mt = Earth magnetic dipole moment = 8.06 x 1022 SI 

Definition of mean (or centred) orbital parameters: 

The centred elements represent the osculating elements for which the short periodic effects from all perturbations 

are removed. They are obtained by filtering the short-period effects of osculating elements which are issued from 

a numerical integration with all the forces (Earth gravity field (a full 70x70 field), Moon and Sun effects, 

atmospheric drag, solar radiation pressure). 

The filtering method uses low-pass filters and the cut-off frequency can be chosen in order to keep the long 

period effects and the secular effects. The filtering method can be applied on keplerian elements as semi-major 

axis, inclination, eccentricity or operational elements of station keeping as altitude, local hour, cross ground tracks 

.... 

Propulsion: 

The constant g 0 is used for hydrazine consumption (∆V = g 0 x Isp x ln(m i/m f)) with g 0 = 9,80665 m/s -2 . 




1.7 REFERENCE AND APPLICABLE DOCUMENTS 

1.7.1 REFERENCE DOCUMENTS 

Reference Document reference Document title 

RD1 LDP.SB.LBP.12.CNES Specification technique de besoin partie 1 plate-forme 

RD2 Issue 5 rev 1 16/06/1999 

PROTEUS Technical Requirements specification 

LDP-SB-LB/LS-12-CNES 

Part 2 : satellite to ground interface 

RD3 LDP-SB-LS-12-CNES PROTEUS specification Technique de Besoin 

Partie 3 : segment sol générique 

RD4 PRO.LB.0.MU.0651.ASC 

issue 1 rev 0 

PROTEUS launch vehicle compatibility guide 

RD5 Deleted 

RD6 PRO.LBP.0.DJ.0640.ASC PROTEUS platform budgets and margins 

RD7 Cours de technologie spatiale CNES - 

Cépaduès éditions 

RD8 Scientific Satellites Achievements and 

Prospects in Europe 20/22/11/96 Paris 

RD9 DGA/T/TI/MS/AM/98022 11/03/98 

Techniques et technologies des véhicules spatiaux (tome 1) 

A new European small platform : PROTEUS, and prospected scientific 

applications 

RD10 PTU/TNT/0034/SE PROTEUS DHU – Interface Data Sheets 

RD11 P-PTU-NOT-0048-SE PROTEUS DHU Analogue Measurement accuracy 

RD12 Guidelines and Assessment Procedures for Limiting Orbital Debris, 

NASA Safety Standard NSS 1740.14 

RD13 JPL D-18663 Rev A, May 20, 2000 Jason 1 Orbital Debris Assessment 




1.7.2 APPLICABLE DOCUMENTS 

None 

1.7.3 STANDARDS 

Reference Document reference Document title 

ST01 MIL-STD-1553 B notice 2 Aircraft Internal Time Division Data Bus 

ST02 MIL-STD-462 Measurement of EMF characteristics 

ST03 ESA PSS-04-106 issue 1 Packet Telemetry Standard 

ST04 ESA PSS-04-107 issue 2 Packet Telecommand Standard 




1.8 ACRONYMS 

1.8.1 REQUIREMENTS ACRONYMS 

The requirements given in the PROTEUS User’s Manual use the following convention 

GR : GRound 

SY : System 

PL : Payload 

xx - yyy - z 

Section number 

GR requirements are applicable to the Ground System 

SY are applicable to all the system that is to say : satellite, payload ... 

PL are applicable to the payload. 

1.8.2 OTHER ACRONYMS 

Requirements number 

AD Applicable Document 

AIT Assembly Integration and Test 

AIV Assembly , Integration and Validation 

AN Analog 

AOCS Attitude and Orbit Control System 

AOS Acquisition Of Signal 

APID Application Process Identifier 

AS16 16-bit Serial Acquisition 

BB Broadband 

BBQ Barbecue 

BC Bus Controller 

BIT Built-in-test 

BDR Baseline Design Review 

BOL Beginning of Life 

BVLE Banc de Validation Logiciel et Electrique (Software and Electrical Validation Bench) 

CCC Command Control Centre 

CCSDS Central Committee for Space Data System 

CDR Critical Design Review 

CLCW Command Link Command Word 

CLTU Command Link Transmission Unit 

CNES Centre National d'Etudes Spatiales 

CoG Centre of Gravity 

CON CNES Operational Network 

COROT COnvection and ROTation 

CS Conducted Susceptibility 

CSS Coarse Sun Sensor 

CST Centre Spatial Toulouse 

CVCM Collected Volatile Condensable Material 

DB Digital Bilevel 

DC Direct Current 

DCN Data Communications Network 

DDV Development Design and Verification 

DHU Data Handling Unit 

DoD Depth of Discharge 




DR Digital Relay 

DS Digital Serial 

EED ElectroExplosive Device 

EEE Electrical, Electronic and Electromechanical 

EGSE Electrical Ground Support Equipment 

EM Engineering Model 

EMC ElectroMagnetic Compatibility 

EMI ElectroMagnetic Interference 

EOL End Of Life 

ESA European Space Agency 

ESD ElectroStatic Discharge 

FDIR Failure Detection Isolation and Recovery 

FDTM Failure Detection TM 

FEM Finite Element Model 

FM Flight Model 

FOV Field of View 

FTP File Transfer Protocol 

GDIS General Design and Interface Specification 

GNSS Global Navigation Satellite System 

GPS Global Positioning System 

GSE Ground Support Equipment 

GYR Gyrometer 

HKTM House Keeping Telemetry 

HKTM-P House Keeping Telemetry Pass 

HKTM-R House Keeping Telemetry Record 

HLC High Level Command 

HW Hardware 

IAT Instrument Aliveness Test 

ICD Interface Control Document 

IDS Interface Data Sheet 

IERS International Earth Rotation Service 

I/F Interface 

IGRF International Geomagnetic Reference Field 

IHCT Instrument Health Check Test 

IIS Instrument Interface Specification 

I/O Input/Output 

IP Internal Protocol 

IPVT Instrument Performance Verification Test 

ISDN Integrated Services Digital Network 

LEO Low Earth Orbit 

LET Linear Energy Transfer 

LGP Local Ground Point 

LISN Line Impedance Stabilised Number 

LISN Line Impedance Stabilised Network 

LNI Local Network Interconnection (CNES Intranet) 

LogB Logbook 

LOS Loss Of Signal 

LTTM Long Term Telemetry 

MAG Magnetometer 

MC Mission Centre 

MCI Mass, Centring & Inertia 

MCR Main Control Room 

MGSE Mechanical Ground Support Equipment 

MIL-STD Military - Standard 




MLI Multi Layer Insulation 

MMIC Microwave Monolithic Integrated Circuit 

MOI Moment Of Inertia 

MSB Most Significant Bit 

MTB Magnetotorquer Bar 

NA Not Applicable 

NB NarrowBand 

NR No requirement 

NOM Normal Operation Mode 

OBSW On Board Software 

OBT On Board Time 

OCM2 Orbit Control Mode 2 Thrusters 

OCM4 Orbit Control Mode 4 Thrusters 

OMP Operations and Manoeuvres Procedures 

OOC Operational Orbit Centre 

OQ Operational Qualification 

OS Operating System 

OVB Operational Validation Bench 

PCE Power Conditioning Equipment 

PDIS Payload Design & Interface Specification 

PDR Preliminary Design Review 

PF Platform 

PFM ProtoFlight Model 

PGGS PROTEUS Generic Ground Segment 

PGR Panel Ground Reference 

PIM Payload Instrument Module 

PL Payload 

PLTM Payload Telemetry 

PM Processor Module 

PPS Pulse Per Second 

PROTEUS Platforme Réutilisable pour l' Observation, les Télécommunications 

et Usages Scientifiques (multimission platform for low Earth orbits) 

PSD Power Spectral Density 

PVT Position / Velocity / Time 

QA Quality Assurance 

QFS Qualification and Flight Spares 

QM Qualification Model 

RAM Random Access Memory 

RD Reference Document 

RDP Rate Damping Phase 

RF Radio Frequency 

RM Reconfiguration Module 

RT Remote Terminal 

RWA Reaction Wheel Assembly 

RX Receiver 

SA Solar Array 

SADM Solar Array Drive Mechanism 

SBDL Standard Balanced Digital Link 

SCC Satellite Control Centre 

SC16 16 bits Serial Command 

SD Satellite Dynamics simulator 

SDB Satellite Data Base 

SGP Single Ground Point 

SHM Safe Hold Mode 




SL Satellite 

SOP Specialised Operations Plan 

SPL Sound Pressure Level 

SPP Sun Pointing Phase 

SR Service Request 

SSGP Standard Control Command Ground System 

STA Star Tracker Assembly 

STAM Star Acquisition Mode 

STR Star Tracker 

SW Software 

TBC To Be Confirmed 

TBD To Be Determined 

TC Telecommand (Ground command) 

TCD Direct Telecommand (hardware TC) 

TCUI Telecommand charge Utile Immédiat (Telecommand Payload Immediat) 

TCUH Telecommand Charge Utile Chargement (Telecommand Payload Software loading) 

TCUT Telecommand Charge Utile « time Tagged » (Telecommand Payload Time Tagged ) 

TQ Technical Qualification 

THR Thrusters 

TM Telemetry 

TMD Direct Telemetry 

TML Total Mass Loss 

TTC Telemetry Tracking and Command 

TTC-ET Telemetry Telecommand Earth Terminal 

TX Transmitter 

UTC Universal Time Coordinated 

VCA Virtual Channel Access 

VCM Virtual Channel Multiplexer 

w/o without 

ZVS Zero Volt Secondaire 








Chapter 2 : Mission envelope 








2.1 LAUNCH VEHICLES 6 

2.1.1 POTENTIAL LAUNCH VEHICLES 6 

2.1.2 LAUNCH VEHICLE ADAPTER 7 

2.2 ACHIEVABLE ORBITS 8 

2.2.1 FLIGHT DOMAIN 8 

2.2.2 CONSTRAINTS RELATIVE TO ORBIT ALTITUDE 9 

2.2.2.1 Environment 9 

2.2.2.2 Global Positioning System (GPS) constraint 9 

2.2.2.3 Attitude and Orbit Control System (AOCS) constraint 9 

2.2.2.4 Telecommunication constraints 9 

2.2.3 CONSTRAINTS RELATIVE TO ORBIT INCLINATION 11 

2.3 IN FLIGHT ORIENTATION AND POINTING 12 

2.3.1 ACHIEVABLE POINTING 12 

2.3.1.1 Earth pointing / fixed yaw / sun synchronous 13 

2.3.1.2 Earth pointing / fixed yaw / Low inclination orbits (About 20 deg) 17 

2.3.1.3 Earth pointing / free yaw / all orbits 18 

2.3.1.4 Inertial pointing 21 

2.3.1.5 Sun pointing 21 

2.3.2 POINTING COMMAND 22 

2.3.3 POINTING AND RESTITUTION PERFORMANCES 22 

2.4 ORBIT DETERMINATION AND CONTROL 23 

2.4.1 ORBIT DETERMINATION PERFORMANCES 23 

2.4.2 ORBIT CONTROL 23 

2.5 FUNDAMENTAL NOTIONS FOR MISSION ANALYSIS 25 

2.5.1 CRITERIA FOR ORBIT DESIGN 26 

2.5.2 DIFFERENT ORBIT TYPES 27 

2.5.2.1 Phased orbits 27 

2.5.2.2 Sun synchronous orbits 28 

2.5.2.3 Frozen orbits 29 

2.5.3 ORBIT PERIOD AND ECLIPSE DURATION 30 

2.5.4 ACCESSIBILITY 32 

2.5.5 VISIBILITY DURATION 33 

2.5.6 ORBITAL PERTURBATIONS 41 

2.5.6.1 Earth potential 41 

2.5.6.2 Moon and Sun gravity potential influence 42 

2.5.6.3 Atmospheric drag 42 

2.5.6.4 Solar radiation pressure 45 

2.5.6.5 Synthesis 45 

2.5.7 ORBITAL MANOEUVRES 46 

2.5.7.1 PROTEUS capabilities 46 

2.5.7.2 Cost in ∆V to change orbital parameters 48 

2.5.7.3 Orbit positioning 50 

2.5.7.4 Orbit maintenance 50 

2.5.7.5 Synthesis 51 

2.5.8 ORBIT DEBRIS GENERATION ANALYSIS 51 





Figure 2.2-1: Orbit envelope .................................................................................................................................. 8 

Figure 2.2-2: Achievable orbits versus launch sites and vehicles ............................................................................ 11 

Figure 2.3-1 : Nominal satellite configuration for the sun synchronous orbits (orbits around noon or midnight on the 

drawing)....................................................................................................................................................... 14 

Figure 2.3-2 : Vertical satellite configuration for the sun synchronous orbits (example : orbits around noon or 

midnight on the drawing).............................................................................................................................. 15 

Figure 2.3-3 : 6 am or 6 pm Sun synchronous orbits............................................................................................. 16 

Figure 2.3-4 : low inclination orbits ...................................................................................................................... 17 

Figure 2.3-5 : Yaw steering .................................................................................................................................. 18 

Figure 2.3-6: Theoretical evolution of the yaw angle along the orbit...................................................................... 19 

Figure 2.3-7: PROTEUS evolution of the yaw angle along the orbit ........................................................................ 19 

Figure 2.3-8: PROTEUS solar array position.......................................................................................................... 20 

Figure 2.5-1: Phased circular orbits 1 to 5 days - the inclination depending on the altitude....................................27 

Figure 2.5-2: Sun synchronous circular orbit inclination versus altitude .................................................................. 28 

Figure 2.5-3 Frozen eccentricity for w = 90° ......................................................................................................... 29 

Figure 2.5-4: keplerian orbital period ................................................................................................................... 30 

Figure 2.5-5: Eclipse duration and percentage of the orbital period depending on the altitude............................... 31 

Figure 2.5-6: Equatorial accessibility of a phased satellite ..................................................................................... 32 

Figure 2.5-7: Station visibility duration (minimum elevation 5°).............................................................................. 34 

Figure 2.5-8: Station vibility duration (minimum elevation 10°) .............................................................................. 34 

Figure 2.5-9 : Kiruna (21.1 E, 67.9 N) (minimum elevation 5°).............................................................................. 35 

Figure 2.5-10: Kiruna (21.1 E, 67.9 N) (minimum elevation 10°)........................................................................... 35 

Figure 2.5-11 : Aussaguel (1.5 E, 43.4 N) (minimum elevation 5°) ........................................................................ 36 

Figure 2.5-12 : Aussaguel (1.5 E, 43.4 N) (minimum elevation 10°) ...................................................................... 37 

Figure 2.5-13 : Kourou (52.6 W, 5.1 N) (minimum elevation 5°) ........................................................................... 38 

Figure 2.5-14 : Kourou (52.6 W, 5.1 N) (minimum elevation 10°) ......................................................................... 38 

Figure 2.5-15 : Hartbeesthoek (27.7 E, 25.9 S) (minimum elevation 5°) ................................................................ 39 

Figure 2.5-16 : Hartbeesthoek (27.7 E, 25.9 S) (minimum elevation 10°) .............................................................. 40 

Figure 2.5-17: Orbital node secular drift, for a circular orbit ................................................................................. 42 

Figure 2.5-18: Solar activity for 01/2003-01/2015 period.................................................................................... 43 

Figure 2.5-19: Maximum DV as a function of the equipped payload mass............................................................. 47 

Figure 2.5-20: Semi major axis correction cost...................................................................................................... 49 

Figure 2.5-21:Inclination correction cost ............................................................................................................... 49 


Table 2.1-1: Main launch vehicles compatible with PROTEUS platform.................................................................... 7 

Table 2.3-1 : PROTEUS satellites pointings............................................................................................................ 12 

Table 2.3-2 : Typical satellite pointing stability ...................................................................................................... 22 

Table 2.4-1 : Platform Inertias in CoG Satellite Reference Frame and CoG position in Satellite Reference Frame.... 24 

Table 2.4-2 : Orbit control performances.............................................................................................................. 24 








LIST OF TABLES...................................................................................................................................................... 3 





LIST OF TBCs 

LIST OF TBDs 

N°§ Sentence Planned Resolution 

§2.2.1 Notice : The lower limit of the flight domain is estimated to 20 deg 

(TBC). 




Chapter 2: Mission envelope 

The first part of this chapter deals with the PROTEUS mission capabilities: the potential launch vehicles, the 

achievable orbits, the possible pointing modes with the associated orbit kinds. The second part is dedicated to the 

User in order to help him analyse his mission at a first level (phase A), choose an orbit type, and assess the required 

propellant capacities in order to achieve the necessary orbital manoeuvres. In any case, the User is strongly 

encouraged to contact either ALCATEL SPACE or CNES in order to detail the mission further on, thus benefiting from 

the greatest experience in mission analysis and design. 

2.1 LAUNCH VEHICLES 

2.1.1 POTENTIAL LAUNCH VEHICLES 

The PROTEUS platform is compatible with the following launch vehicles: Ariane 5, Athena 2, Cosmos, Delta 2, LM- 

2D, PSLV, Rockot, Soyuz and Taurus. Their main characteristics are summarised in Table 2.1-1, for information only. 

The User shall refer to the corresponding launcher manual applicable at its study moment 

This list takes into account all developed launch vehicles in the 500 to 1000 kg class. It could be updated if new 

launch vehicles became available. Assuming this large panel of compatible launch vehicles, it will be easy to adapt 

PROTEUS with other launch vehicles like Ariane 4 (Europe), Atlas (USA), etc... 




Launch 

vehicle 

Country / 

Launch sites 

Launch 

service 

provider 

First 

flight 

Usable volume 

diameter (mm) 



Comments 

Ariane 5 Europe/Kourou Arianespace 1998 4570 or 4800 Multiple launch 

Athena 2 USA/Cape Lockeed January 1984 

(LMLV2) Canaveral, 

Vandenberg 

Martin 1998 

Cosmos Russia/Plesetsk Cosmos 

international 

1970 2200 

Delta 2 USA/Cape Boeing 1995 2743 (upper Dual launch 

(upper Canaveral, 

position) 

position) Vandenberg 

2330 (lower 

position) 

LM-2D China/Jiquan GWIC 1992 2360/1715 

PSLV India/Shriarikota ISRO 1993 2900 

Rockot Russia/Plesetsk Eurockot 1990 1983 Dual launch 

foreseen 

Soyuz Russia/Baikonur, Starsem in the 3395 

Plesetsk 

sixties 

Taurus USA/Cape 

OSC 1994 2055 Taurus versions 

Canaveral, 

equipped with a 

Vandenberg, Wallops 

2.34 m fairing * 

Table 2.1-1: Main launch vehicles compatible with PROTEUS platform 

* The Taurus, Taurus XL, Taurus XLS can be equipped with a 2.34 m diameter fairing. 

2.1.2 LAUNCH VEHICLE ADAPTER 

The standard PROTEUS bottom frame has a standard interface with 60 M5 screws on a 943,6 mm diameter. 

The launch vehicle adapter is the interface hardware between the platform bottom frame and the launch vehicle. It is 

bolted on the platform and is maintained by a clamp band on the launch vehicle side. At the launch vehicle/satellite 

separation, the clamp band opens and the satellite (with the launch vehicle adapter staying fixed at the platform) 

separates off from the launch vehicle (cf. chapter 1.5). The launch vehicle adapter is a thick ring with a 943.5 mm 

diameter (for Delta 2, Taurus, Athena 2); but it can be different depending on the launch vehicle (interface diameter). 

ALCATEL SPACE and CNES are logically responsible for the launch vehicle adapter mechanical and thermal design; 

it shall be negotiated with the launch vehicle authorities case by case.


2.2 ACHIEVABLE ORBITS 

2.2.1 FLIGHT DOMAIN 

The allowed orbits are limited by several constraints detailed hereafter. The resulting orbit envelope is shown on 

Figure 2.2-1. 

Altitude 

2000 Km 

1750 Km 

1500 Km AOCS 

Limit. 

1250 Km 

20° (TBC) 

1000 Km 

750 Km 

0 Km 

PROTEUS FLIGHT ENVELOPE 

Allowed with life duration (radiations) and other 

restrictions at upper altitudes (GPS, magnetic field,...) 

500 Km 

Allowed with life duration (monoatomic oxygen, atmos. drag) 

250 Kmand 

ground station visibility restrictions 

0° 20° 40° 60° 80° 100° 120° 140° 

Inclination 

Figure 2.2-1: Orbit envelope 

Notice : The lower limit of the flight domain is estimated to 20 deg (TBC). 



3 Years w ith margin 

(5 Years Without margins) 

3 Years w ithout margins 

Allowed with 

launch site 

restrictions 

(Sun synchronous orbits)


2.2.2 CONSTRAINTS RELATIVE TO ORBIT ALTITUDE 

2.2.2.1 Environment 

The lower altitudes limit of the flight domain is determined by the atmospheric drag and the mono-oxygen effects. 

The atmospheric drag is usually compensated by periodic manoeuvres in order to maintain the altitude and/or the 

semi-major axis or the orbit. In this area, the propellant capacity of the satellite limits the orbital manoeuvres 

capacity and has a direct effect on the satellite lifetime. 

The mono-atomicoxygen contained in the upper atmosphere reacts with satellite materials, especially Kapton and 

Silver and causes the erosion and the weakening of these materials. 

MLI external layer is very sensitive to mono-atomic axygene dose. Standard PROTEUS design allows minimum 

altitude of about 600 km. 

The mono-atomic oxygene dose specified for the Proteus generic Star Tracker is such that STR is not designed for 

missions at an altitude under 600 km. 

Standard Solar Arrays (without protection against atomic oxygen) can stand environment met at altitude around 600 

km. With a coating protection on the solar arrays, the flight domain can cover lower altitudes. 

Due to the exponential relationship between these effects and the altitude, a minimum altitude of 600 km is 

recommended. 

The upper altitudes (above 1500km, roughly) are sensitive to radiation, LET (Linear Energy Transfer) and Trapped 

Proton fluxes. 

The sizing has been performed taking into account a cumulated radiation dose over 5 years on a reference orbit : 

1336 km/66° (without margins), cf. red curve on figure 2.2-1. 

2.2.2.2 Global Positioning System (GPS) constraint 

A GPS is used on board to have a time, position and velocity reference. 

High altitudes limit the GPS constellation satellites visibility, but GPS satellites are far above the satellites using the 

PROTEUS platform: 20 000 km versus about 1500 km and therefore are not a limiting constraint for PROTEUS 

nominal flight envelope. 

A specific study (on request/orbit dependent) will be done for circular orbits or elliptical orbits with an apogee higher 

than 1500 km or in case of inertial pointing (mission dependent). 

2.2.2.3 Attitude and Orbit Control System (AOCS) constraint 

The PROTEUS platform AOCS uses magnetic torquers for reaction wheels desaturation in normal mode and attitude 

acquisition in Safe Hold Mode. This equipment can generate a torque perpendicular to the Earth magnetic field. 

Unfortunately, at the equator, the Earth’s magnetic field is nearly perpendicular to the equatorial plane, so one axis 

becomes poorly controllable for low inclination orbits. Below 20° inclination and depending on satellite inertia, the 

mission feasibility has to be checked on a case by case basis. 

The magnetic field strength decreases with altitude and an other limitation could appear for very high orbits. As for 

GPS constraint, a specific study will be led for circular orbits or elliptical orbits with an apogee higher than 1500 km. 

2.2.2.4 Telecommunication constraints 

The downlink and ground command budget has been evaluated for a circular orbit with a 1336 km altitude and a 

minimum elevation angle equal to 10 deg, which is nearly the highest altitude allowed. 




For higher orbits and the same site angle, the satellite to ground station distance is greater, but the visibility duration 

per day with a single ground station increases. RF link budget performances could be maintained assuming a higher 

minimum elevation angle and/or a data rate reduction. 

For lower altitudes, the visibility duration per day decreases dramatically and a second ground station could be 

necessary (mission dependent on TM flow and ground station minimum elevation constraint) 




2.2.3 CONSTRAINTS RELATIVE TO ORBIT INCLINATION 

The achievable orbits are limited by the launch azimuth allowable from a given launch site. PROTEUS can correct the 

launch vehicle injection errors, but manoeuvres to change significantly the orbit inclination are propellant 

consuming. The achievable orbits depending on the launch vehicles and launch sites are shown on Figure 2.2-2. 

Figure 2.2-2: Achievable orbits versus launch sites and vehicles 




2.3 IN FLIGHT ORIENTATION AND POINTING 

2.3.1 ACHIEVABLE POINTING 

The standard PROTEUS platform allows five main kinds of pointing: 

Earth pointing with a fixed yaw on a sun synchronous orbit or on a low inclination orbit (about 20 deg), 

Earth pointing with a yaw steering on every orbit, 

inertial pointing, 

sun pointing on a sun synchronous orbit, 

other non standard pointing modes which can be studied upon request. 

The possible manoeuvres around these pointings shall be compatible with the platform reaction wheels torque 

capacity and the moment of inertia acceptable for the reaction wheels. 

The AOCS dynamic range is compatible with the following pointings : Earth pointing, Nadir pointing, track pointing, 

yaw steering, inertial pointing. 

Table 2.3-1 summarises the conceivable satellite pointings considering the PROTEUS flight domain. 

Table 2.3-1 : PROTEUS satellites pointings 

The pointing chosen according to the mission needs imposes: 

the satellite orientation in routine mode, so the associated mechanical configuration of the payload and the set 

up of some equipment components such as the star trackers, the antennas. 

(satellite lay out, centering, inertia) 

thermal limitations for the satellite, 

solar arrays efficiency and so the power limitations for the satellite and the payload, 

the AOCS approach, 

telecommunications link constraints. 

For PROTEUS based missions, these constraints lead to consider preferably the standard satellite configurations 

described hereafter. 




2.3.1.1 Earth pointing / fixed yaw / sun synchronous 

The satellite configurations presented in this paragraph can be applied for fixed yaw but also for angular variations 

of a few degrees around the pointing axis. 

2.3.1.1.1 Sun synchronous orbits 

2.3.1.1.1.1 Definition 

The orbital plane keeps a constant angle with the Earth-Sun direction during the year. Under some inclination and 

altitude conditions, the orbital plane drift is equal to the Earth movement around the Sun (0.986 deg per day). 

Sun synchronous orbits allows to get: 

a constant solar local time at a reference location, which determines a constant illumination (if the seasonal 

variations are not considered), 

a sweeping over the whole surface of the Earth; the orbit is nearly polar (orbit inclination around 98 deg). 

They are usually circular with the frozen perigee and at a constant altitude. This orbit kind is typically selected for 

Earth observation. 

2.3.1.1.1.2 Satellite pointing 

For PROTEUS, the possible sun synchronous orbits are as follows: 

sun synchronous orbits with an ascending or a descending node between 9:30 am and 2:30 pm or between 

9:30 pm and 2:30 am. 

sun synchronous orbits with an ascending or a descending node close to 6 am or to 6 pm. In this case, the 

nominal satellite configuration is classical. 

For the sun synchronous orbits, two satellite pointing modes are possible: 

the +Zs axis can be oriented towards the Earth; it is called « nominal satellite configuration » 

the +Xs axis can be oriented towards the Earth; it is called « vertical satellite configuration ». This second 

configuration shall be negotiated. Currently, the main identified critical points for this configuration are : 

TMTC link 

Thermal control (mainly battery and DHU) 

Lead to important attitude slew for orbit manoeuvers and loss of pointing 

GPS antenna field of view (only on –Z s side and –X s sides) 

STA accommodation 




Sun synchronous orbits with an ascending or a descending node between 9:30 am and 2:30 pm or between 9:30 

pm and 2:30 am 

a) Nominal configuration 

The nominal satellite configuration (cf. Figure 2.3-1) is such that the satellite +Zs axis points towards the Earth. 

The +Ys axis is aligned with the solar array rotation axis and it is oriented such that the Sun is in the - Ys 

hemisphere. The +Xs axis is aligned with the launch vehicle axis and is oriented following the velocity or antivelocity 

direction depending on local time. The axis direction is imposed by the right handed orthogonal 

reference frame. 

+X S 

+Z S 

+Y S 

Figure 2.3-1 : Nominal satellite configuration for the sun synchronous orbits (orbits around noon or 

midnight on the drawing) 




b) Vertical configuration 

For this orbit kind, the payload accommodation with the platform is such that the satellite adopts the vertical 

configuration (cf. Figure 2.3-2); that means the +Xs axis is pointed towards the Earth. The +Ys axis is aligned 

with the solar array rotation axis and it is oriented such that the Sun is in the - Ys hemisphere. The +Z axis is 

aligned with the launch vehicle axis and is oriented following the velocity or anti-velocity direction depending 

on local time. The axis direction is imposed by the right handed orthogonal reference frame. 

+X S 

+Z S 

Figure 2.3-2 : Vertical satellite configuration for the sun synchronous orbits (example : orbits around 

noon or midnight on the drawing) 

+Y S 




Sun synchronous orbits with an ascending or a descending node from 5 am to 7 am or 5 pm to 7 pm. 

In this case, the Sun/orbital plane angle is equal to the orbit inclination (nearly 98 deg), plus a seasonal variation 

equal to the solar declination (+23.5 deg maximum at solstice). 

The satellite will fly with the +Ys axis parallel to the orbital speed and preferably the Sun is in - Xs hemisphere to 

have a configuration similar to the Safe Hold Mode pointing. 

+Y S 

+Z S 

+X S 

Figure 2.3-3 : 6 am or 6 pm Sun synchronous orbits 




2.3.1.2 Earth pointing / fixed yaw / Low inclination orbits (About 20 deg) 

Low inclination orbits drift around the polar axis (several degrees a day) and the sun/orbit plane angle varies in the 

following interval [-(orbit inclination + solar declination);+(orbit inclination + solar declination)]. It is necessary to 

make a 180° slew around the yaw axis (Zs) when the sun crosses the orbital plane to maintain the sun in one satellite 

hemisphere. A three axis pointed satellite flies with the +Xs axis along the orbital velocity during half the time and 

with the +Xs axis in the opposite direction other wise. 

+Y S 

+Z S 



+X S 

Figure 2.3-4 : low inclination orbits


2.3.1.3 Earth pointing / free yaw / all orbits 

For Earth pointing, the +Zs satellite axis remains pointed towards the Nadir axis. In the free yaw case, the satellite 

attitude follows a yaw steering to orient the solar arrays towards the sun (cf. Figure 2.3-5); the aim is to avoid 

thermal and power losses. 

The satellite rotates around the Earth direction to maintain the Sun in the (Xs, Zs) plane, with the Sun in the -Xs 

hemisphere (configuration close to the Safe Holdmode). Then the solar arrays are continuously oriented 

towards the Sun, following a near sinusoidal movement along the orbital period. 

When the Sun angle versus the orbital plane is less than 20 deg (typical value), the yaw steering movement is 

stopped and the satellite follows a three axis pointing profile with +Xs or -Xs oriented towards the orbital 

speed. 

Figure 2.3-5 : Yaw steering 

The yaw angle theoretical evolution versus Sun/orbital plane is shown on Figure 2.3-6. 

Figure 2.3-7 and Figure 2.3-8 show the yaw angle and the solar array position given by the implemented 

approximated laws. 



.. 

.. 


Yaw angle (deg) 

Yaw angle (deg) 

180 

160 

140 

120 

100 

80 

60 

40 

20 

0 

-20 

-40 

-60 

-80 

-100 

-120 

-140 

-160 

-180 

0 

180.0 

160.0 

140.0 

120.0 

100.0 

80.0 

60.0 

40.0 

20.0 

0.0 

-20.0 

-40.0 

-60.0 

-80.0 

-100.0 

-120.0 

-140.0 

-160.0 

-180.0 

20 

40 

60 

80 

100 

120 

140 

160 

180 

200 



220 

240 

260 

280 

Orbital position with regard to subSun position (deg) 

Figure 2.3-6: Theoretical evolution of the yaw angle along the orbit 

0 

10 

20 

30 

40 

50 

60 

70 

80 

90 

100 

110 

120 

130 

140 

150 

160 

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180 

190 

200 

210 

220 

230 

240 

250 

260 

270 

280 

290 

300 

310 

320 

330 

340 

350 

360 


Figure 2.3-7: PROTEUS evolution of the yaw angle along the orbit 

300 

320 

340 

360 

Beta=-90° 

Beta=-75° 

Beta=-60° 

Beta=-45° 

Beta=-30° 

Beta=-15° 

Beta=90° 

Beta=75° 

Beta=60° 

Beta=45° 

Beta=30° 

Beta=15° 

Beta -90 deg 

Beta -75 deg 

Beta -60 deg 

Beta -45 deg 

Beta -30 deg 

Beta -15 deg 

Beta 15 deg 

Beta 30 deg 

Beta 45 deg 

Beta 60 deg 

Beta 75 deg 

Beta 90 deg


Solar array position (deg) 

90.0 

80.0 

70.0 

60.0 

50.0 

40.0 

30.0 

20.0 

10.0 

0.0 

-10.0 

-20.0 

-30.0 

-40.0 

-50.0 

-60.0 

-70.0 

-80.0 

-90.0 

0 

10 

20 

30 

40 

50 

60 

70 

80 

90 

100 

110 

120 

130 

140 

150 

160 

170 

180 

190 

200 

210 

220 

230 

240 

250 

260 

270 

280 

290 

300 

310 

320 

330 

340 

350 

360 


Figure 2.3-8: PROTEUS solar array position 



Beta +/-15 deg 





Beta +/-90 deg


2.3.1.4 Inertial pointing 

In order to avoid Earth shadowing, the polar inertial orbit is recommended. 

For inertial missions, the payload will have its field of view boresight towards +X satellite axis. 

It is supposed that these missions do not impose attitude around X axis (inertial, but 2 axis pointing) and the Sun 

could remain in the -X satellite hemisphere. Then, the attitude around the X axis will be chosen such that the Sun is 

near (X,Z) plane to minimize thermal constraints and optimize the power budget: 

the Sun could be placed perpendicular to the Solar Array by a rotation of these Solar Arrays around these axis. 

the Sun could remain in the +Z hemisphere to minimize solar aspect on the battery radiator, performing a 

180° slew around +X when the Sun crosses the (X,Y) plane (solar aspect up to 10 to 20° maximum on the -Z 

face tolerated). 

Star trackers orientation will be optimized (near payload boresight) to avoid Earth, Sun, Moon, planets, stars holes 

perturbations. 

These limitations can be reviewed on a case by case analysis due to the difficulty to define a generic inertial pointing 

mission and associated constraints. 

2.3.1.5 Sun pointing 

Sun pointing can be considered as a particular case of inertial pointing, with the Sun along the +Xs or -Xs direction. 

In order to avoid Earth shadowing (Sun eclipse), a Sun synchronous orbit with a 6 am or 6 pm node and with a high 

altitude is recommended. The pointing direction (Xs axis) is oriented between 0 and 32 deg from the perpendicular 

to the orbit. 




2.3.2 POINTING COMMAND 

The pointing command is defined by time tagged commands. These commands are defined by: 

the time T 0 from which the time tagged command shall be applied 

the attitude quaternion at T 0 ( the command shall be continuous with the preceding command) 

the quaternion evolution rate from T 0 . This evolution rate is described by: 

a flag indicating if the evolution is versus an Inertial Frame or the Local Orbital Frame. 

a flag indicating if the evolution is described by a polynome defined versus T 0 or a Fourier serie defined 

versus the orbital position ωt. 

the degree of the polynomial or the Fourier Serie (maximum value 4). 

the polynomial or Fourier serie coefficients along the three axes. 

the Solar array commanded evolution. This command is defined as the sum of a linear function and a Fourier 

serie of degre 1: 

a0+a1t+b1sin ωt +c1 cos ωt 

These commands are applied until a new time tagged command replaces them (infinite duration possible). 

For instance, a geocentric mission has a command with all coefficients at 0, so this command is valid for the whole 

mission duration. 

2.3.3 POINTING AND RESTITUTION PERFORMANCES 

The pointing requirement is mission dependent. It consists in two main items : 

Alignment difference (bias, thermoelastic…) between payload boresight and STA interface plane (mission and 

payload dependent) 

Pointing performance at STA interface plane level with respect to reference frame (inertial or local orbital 

frame) 

The satellite system provides this pointing performance at the STA interface plane with an accuracy of 0,05 deg (3σ) 

around each axis. 

In order to achieve this performance, the payload shall fulfil the two requirements given in section 3.1.4.3. 

The platform pointing stability in routine is mission dependent. 

Without perturbation due to the payload, the typical values are indicated in Table 2.3-2. 

Frequency band pointing stability (3σ) 

0.1 to 1 Hz 7.10 -4 deg/s 

1 to 5 Hz 3.10 -4 deg 

5 to 20 Hz 10 -2 deg/s 

20 to 80 Hz 2.10 -4 deg 

>80 Hz 3.10 -2 deg/s 

Table 2.3-2 : Typical satellite pointing stability 




2.4 ORBIT DETERMINATION AND CONTROL 

2.4.1 ORBIT DETERMINATION PERFORMANCES 

The measurements to determine the orbit are performed with the on board GPS. There are three methods to 

determine the orbit: 

real time on board measurements: the orbit restitution accuracy is estimated to 120 m (3σ), 

post-processing of the telemetry data: in this case, the orbit restitution accuracy is better than 30 m (3σ) 

(this value must be confirmed for the case of very low orbit with an altitude < 800 km and depends 

on solar activity) 

the orbit prediction depending on the initial error issued from the orbit restitution and the error issued from the 

orbit perturbation. In the low orbit case, the atmospheric drag effect is estimated with difficulties, affecting the 

orbit prediction accuracy. The orbit prediction accuracy is along the satellite track and strongly depends on the 

orbit altitude and on the solar activity. 

2.4.2 ORBIT CONTROL 

The PROTEUS platform is equipped with an hydrazine propulsion system which allows 

a complementary injection after the launch phase, 

to acquire the orbit with accuracy and to correct for launch errors, 

to maintain the orbit. 

The orbital manoeuvres resulting from these three operation types must correspond to an overall velocity increment 

∆V equal to : 

∆V = 130 m/s (for the 450 kg satellite class). 

The detail for other satellites is given in chapter 2.5.7. 

The possible minimal magnitude of a manoeuvre is estimated to 0,5 mm/s, and the maximal one is equal to 5 m/s. 

The manoeuvres resolution is better than 0,5 to 1 mm/s depending on the number of thrusters used (2 to 4 which 

corresponds to the OCM2 and OCM4 modes). The accuracy to perform the manoeuvres is better than 5% after the 

in flight calibration. 

The delay between two orbital corrections is typically one orbit duration for inclination corrections and 0.5 orbit 

duration for semi major axis corrections. The delay between two manoeuvres depends on the visibility characteristics; 

during the first visibility, a telemetry allows to know with accuracy the orbit just after the first manoeuvre and during 

the second visibility, a telecommand is sent to perform the next manoeuvre. 

The satellite slew rate depends on the inertia of the payload and of the wheels torque capacity. In the nominal case, 

the available torque is 0.1 Nm (worst case) on each axis for a maximum duration of nearly 1 minute. 

For information, Platform inertias and Center of Gravity (CoG) position are given in satellite co_ordinate system with 

solar arrays folded and unfolded (cf. Table 2.4-1). These platform characteristics do not take into account STA and 

launch vehicle adapter characteristics. 




Folded configuration Unfolded configuration 

Platform Inertias 

Ix (m2 *kg) 55 425 

Iy (m2 *kg) 45 45 

Iz (m2 *kg) 

Platform CoG position 

45 415 

X (mm) 480 525 

Y (mm) 0 0 

Z (mm) -10 -10 

Table 2.4-1 : Platform Inertias in CoG Satellite Reference Frame and CoG position in Satellite 

Reference Frame 

The mission interruption duration depends on the flight satellite configuration : 

For a standard flight configuration with the +Xs satellite axis aligned with the velocity direction in the Nadir or 

Earth pointing case, orbital manoeuvres shall not imply any flight configuration change, so the payload should 

stay pointed to the same direction. In this case, orbital manoeuvres should not imply any mission interruption. 

But during these manoeuvres, the pointing performance could be damaged. 

For a vertical flight with the Zs satellite axis aligned with the velocity direction, the satellite shall be rotated by 

90° before performing orbital manoeuvres. It will imply mission interruption; the duration will depend on 

satellite inertia and the time to perform the manoeuvres. 

All these main performances are summarised in Table 2.4-2. 

Characteristic values 

overall velocity increment ∆V 130 m/s (for the 450 kg satellite class) 

(detail given in chapter 2.5.7) 

manoeuvres magnitude : 

minimal 

maximal 

1 mms 

2.5 m/s 

Manoeuvres resolution 0.5 - 1 mm/s 

Manoeuvres accuracy 5% 

Delay between two orbital corrections : 

for inclination corrections 

for semi- major axis corrections 

1 orbit duration 

0.5 orbit duration 

Available torque on each satellite axis 0.1 Nm for about 1 minute 

Table 2.4-2 : Orbit control performances 




2.5 FUNDAMENTAL NOTIONS FOR MISSION ANALYSIS 

This chapter is a general introduction to the spatial mechanics for low Earth orbits. These first rules allows the User to 

perform his mission analysis, they are not specific to PROTEUS based missions. 

The user must choose the orbit kind fulfilling the technical and scientific needs of his mission. At first, the criteria 

necessary to select the orbit are listed; the aim is to decide on the main orbital parameters without impacting the 

platform performances, without restricting the specifications at the payload level and to achieve the mission 

objectives. 

Some useful notions for the orbit choice are reminded : 

the main orbit kinds are listed with an abacus which allows to determine the orbit plane position for every 

case; this may imply an orbit inclination depending on the altitude, 

the orbit period and the eclipse duration depending on the altitude are given in the keplerian orbit case, 

the main forces which can disturb the satellite motion and their impact are described, for instance the daily 

inclination drift and the altitude drift versus time. 

the visibility duration depending on the altitude, the elevation, the different stations used. 

These notions are applied for circular orbits at an altitude between 400 and 1000 km and with an inclination 

between 0 and 100 deg because these orbits are considered as the most usual ones for PROTEUS missions. Highly 

elliptical or other particular orbits are conceivable but need specific studies. 

The second part of this chapter deals with the orbital manoeuvres. It allows to determine the cost in ∆V (and 

therefore in propellant mass) so that the satellite is able to reach and to maintain the chosen operational orbit even 

when launch errors and orbital perturbations are considered. Then, the User can estimate if PROTEUS is able to fulfil 

the mission according to the chosen orbit. 




2.5.1 CRITERIA FOR ORBIT DESIGN 

The orbit choice criteria are classified into two groups: the first one concerns items having a direct impact on the 

platform performances, the second list only concerns the payload. In the latter, the most usual items are reported 

assuming each payload has very specific needs. 

Main criteria which have an impact on the platform performances: 

the pointing type, the payload accommodation and the orbit are coupled (see paragraph 2.3), 

over 1000 km in altitude, the radiation effects become important and affect the mission life duration, 

under 600 km in altitude, the atmospheric drag implies a propellant consumption increase to maintain the 

orbit; mono-oxygen reacts with satellite external materials like Kapton or solar cell connections, meaning that 

mission duration is affected, 

the ground visibility duration with one Earth terminal leads to select a high inclination orbit, or an equatorial 

orbit and a high altitude, 

the launch vehicle cannot reach all inclinations because of the launch pad latitude and/or azimuth restrictions. 

As a satellite can not procure a high orbit inclination modification on its own, the orbit inclination choice does 

not only depend on the payload needs; the launch pad location and the launch vehicle performances also 

have to be considered. For instance, an orbit inclination lower than 28.5 deg cannot be reached with a small 

launch vehicle from its usual launch pad, without an important dogleg (change of orbital plane by the launch 

vehicle needs an important propellant consumption) 

the link budget optimisation leads to minimise the altitude for radar or telecommunication missions. 

Main criteria for orbit design not impacting the platform performances: 

the resolution for an optical mission leads to choose a low altitude for the orbit, 

the accessibility of the mission leads to increase the altitude and constrains some particular altitudes 

depending on the mission needs, 

the altitude repetitivity from one orbit to another one leads to select a frozen eccentricity. 




2.5.2 DIFFERENT ORBIT TYPES 

Hereafter are described some particular orbits: 

phased orbits for which the satellite flies over the same ground track with a periodical cycle, 

sun synchronous orbits for which the orbital plane keeps a constant angle with the Earth-Sun direction, 

frozen orbits for which the perigee argument and the eccentricity are constant. 

2.5.2.1 Phased orbits 

A phased orbit ensures the periodicity of the satellite ground track. That means the satellite performs a daily 

revolution number corresponding to a rational fraction p = n+m/q with n = number of full revolutions per day, m = 

sub cycle duration (in days), q = cycle duration (in days). In this case, the ground track will have a period of q days. 

Chart 2.5.1 gives the inclination depending on the circular phased orbit altitude. The orbital perturbation taken into 

account is the J 2 effect for Earth gravitational potential, due to the Earth oblateness. 

Figure 2.5-1: Phased circular orbits 1 to 5 days - the inclination depending on the altitude 




2.5.2.2 Sun synchronous orbits 

For Sun synchronous orbits, the orbital ascending node drift is equal to the Sun mean apparent rate 

Ω 1 = n S = 0.985626 deg/day. Figure 2.5-2 shows the inclination for circular Sun synchronous orbits in the typical 

altitude range for PROTEUS based missions. 

Figure 2.5-2: Sun synchronous circular orbit inclination versus altitude 




2.5.2.3 Frozen orbits 

Frozen orbits are chosen in certain missions to obtain a repetitive altitude at any given latitude. This is used for radar 

applications, where the range needs to be determined with accuracy. In order to maintain a constant altitude, the 

secular variations of average parameters such as the eccentricity e and the argument of perigee ω which are under 

Earth potential perturbation effects must be cancelled. 

The following equation set is solved: de/dt =f1(e, ω,u) = 0 and dω/dt = f2(e, ω,u)=0, where u corresponds to the 

orbital parameters such as the semi major axis a and the inclination i, f1 and f2 being temporal functions linked to 

the Earth potential. 

Assumption: the expression used for the Earth potential to estimate the solutions includes the secular and long period 

terms (the zonal terms, up to degree 50), short period perturbations being considered to be negligible. 

There are two optimum values of the frozen perigee ω G = 90° and ω G = 270°. Figure 2.5-3 shows the frozen 

eccentricity depending on the couple of parameters (a, i). 

Figure 2.5-3 Frozen eccentricity for w = 90° 

The eccentricities given on the Figure 2.5-3 are to be multiplied by 10 -3 . 




2.5.3 ORBIT PERIOD AND ECLIPSE DURATION 

The objective is to give a rough idea of the orbit period and eclipse duration. These values are estimated under the 

keplerian orbit assumption. This model means that the only force applied to the satellite is the central force expressed 

in 1/r2 and caused by the Earth gravity. The keplerian period depends on the altitude for circular orbits as shown on 

Figure 2.5-4. The maximum eclipse duration depends on the altitude, and appears on Figure 2.5-5. 

.. 

Figure 2.5-4: keplerian orbital period 




Figure 2.5-5: Eclipse duration and percentage of the orbital period depending on the altitude 




2.5.4 ACCESSIBILITY 

This property assesses the capacity to access the satellite from a given location on Earth. In the case of phased orbits 

for instance, it is very useful to evaluate the duration of the cycle for full access of the satellite to the equator over its 

orbital cycle. Figure 2.5-6 shows, as a function of altitude and for various orbital cycle durations, the half field of 

view of the instrument (or antenna) required for a full equatorial coverage. A calculation of this kind is necessary for 

each possible orbit, in order to estimate the performance of the mission. 

Figure 2.5-6: Equatorial accessibility of a phased satellite 




2.5.5 VISIBILITY DURATION 

The PROTEUS based satellite - ground link is ensured by a S-band TM/TC link which characteristics are detailed in 

chapter 9. Except for vertical flight (some restrictions to be analysed on a case by case basis), PROTEUS allows a 

TM/TC link budget with a margin greater than 3 dB for all altitudes from 400 to 1800 km and for all elevations over 

5 deg (with no mask for ground antenna). For a circular orbit, the visibility duration of PROTEUS only depends on the 

orbit altitude, minimum elevation, and station localisation versus orbital track. 

The control station visibility typical duration is estimated to 10 minutes in the case of low orbits characterised by a 

period of around 100 minutes. 

The User can estimate the visibility duration for his mission with the following graphs. Depending on the ground 

station location, the visibility duration may be reduced due to some geometrical masks for elevation between 5 deg 

and 10 deg. Before the choice of the ground station location, the visibility duration budget shall be done with an 

assumption of a minimum elevation of 10°. 

Figure 2.5-7 gives the station visibility duration depending on the maximum elevation when the satellite enters the 

visibility area (defined by a minimum 5 elevation, here), for low orbits (altitudes between 400 km and 1800 km). 

Figure 2.5-8 gives the same think for a minimum 10° elevation. 

A computation of the duration of the accesses of the satellite to the associated ground station is necessary for each 

specific mission, in order to make sure that the link budget is adapted to the downloading needs of the mission, 

given the capabilities of the TM/TC function. 

For information, Figure 2.5-9, 2.5-11, 2.5-13 and Figure 2.5-15 give for Kiruna (Sweden), Aussaguel (France), 

Kourou (French Guiana) and Hartbeesthock (HBK South Africa) stations the mean daily visibility duration function of 

altitude and inclination with a minimal elevation equal to 5°. Figure 2.5-10, 2.5-12, 2.5-14 and 2.5-16 give the 

same think for a minimal elevation equal to 10°. 



.. 

.. 


Figure 2.5-7: Station visibility duration (minimum elevation 5°) 

Figure 2.5-8: Station vibility duration (minimum elevation 10°) 



.. 

.. 


Figure 2.5-9 : Kiruna (21.1 E, 67.9 N) (minimum elevation 5°) 

Figure 2.5-10: Kiruna (21.1 E, 67.9 N) (minimum elevation 10°) 



.. 


Figure 2.5-11 : Aussaguel (1.5 E, 43.4 N) (minimum elevation 5°) 



.. 


Figure 2.5-12 : Aussaguel (1.5 E, 43.4 N) (minimum elevation 10°) 



.. 

.. 


Figure 2.5-13 : Kourou (52.6 W, 5.1 N) (minimum elevation 5°) 

Figure 2.5-14 : Kourou (52.6 W, 5.1 N) (minimum elevation 10°) 



.. 


Figure 2.5-15 : Hartbeesthoek (27.7 E, 25.9 S) (minimum elevation 5°) 




Figure 2.5-16 : Hartbeesthoek (27.7 E, 25.9 S) (minimum elevation 10°) 




2.5.6 ORBITAL PERTURBATIONS 

In the keplerian orbit case, the satellite only experiences a central force behaving in 1/r2 . In reality, the satellite 

movement is perturbed by other forces: 

gravitational disturbances: oblateness of the Earth, Sun and Moon gravitational attraction, 

non gravitational disturbances: solar radiation pressure, atmospheric drag. 

2.5.6.1 Earth potential 

Earth potential is neither spherical nor homogeneous. It can be described as the sum of spherical harmonics. C l,m 

and S l,m are the harmonic coefficients with the degree l and the order m. The spherical harmonics can be classified 

into two groups: 

the zonal harmonics (m = 0, J n = –C no) correspond to the irregularities in latitude 

the tesseral harmonics (m ≠ 0, m ≠ l) corresponds to the irregularities in longitude 

The first zonal harmonic (J 2) is about 10 -3 in comparison with the main term in µ/r whereas higher terms have a 

magnitude lower or equal to 10 -6 . In order to have better results, the model for Earth potential often takes into 

account the J 2 harmonic. 

According to this assumption, the secular variations of orbital elements of the orbital node Ω and the perigee 

argument ω spell in the following way: 

the orbital node drift in deg/day: 

the perigee argument drift in deg/day: 

Ω 

1 

1 

Ω 

1 

= 

3 

ω = − 

1 

2 

3nJ 

2Rt 

cosi 

= − 2 2 2 

2( 

1− 

e ) a 

− 2. 

064 

x 10 

2 2 7/2 ( 1- 

e ) a 

2 

nJ 2Rt 

1. 

032 

ω = − 



14 

cos i 

( 1− 

5cos 

2 2 2 

4( 

1− 

e ) a 

2 

i) 

14 

2 

x 10 ( 1− 

5 cos i) 

2 2 7/2 ( 1− 

e ) a 

with n = Keplerian average angular velocity, a= semi major axis in km, i = inclination in deg. 

Figure 2.5-17 shows the orbital node drift depending on the (inclination, altitude) couple for a circular orbit.


Figure 2.5-17: Orbital node secular drift, for a circular orbit 

2.5.6.2 Moon and Sun gravity potential influence 

As the Moon is attracted by the Earth and the Earth by the Sun, a satellite in space is attracted by more than one 

celestial body. There are usually long period effects (1 year) on the inclination i, the orbital node Ω and the position 

on orbit α and on the secular terms on Ω. In the Sun synchronous orbit case, there is a secular drift on i and Ω. 

The secular drift due to the effects of the Sun and Moon on i and Ω strongly depends on the local hour of the 

ascending node: this property can be used to reduce this drift value. 

2.5.6.3 Atmospheric drag 

Whereas the previous perturbations have a gravitational origin, the atmospheric drag creates areal forces. 

The aerodynamic force F, caused by the atmospheric drag up to a 1000 km altitude, gives the corresponding 

acceleration: 

1 2 S 

γ = − ρV 

CX 

µ 

2 m 

 

where ρ is the atmospheric density, V the spacecraft velocity versus the atmosphere, S the drag effective surface, CX the drag coefficient, m the mass (Cx depends on the altitude; Cx (400 km) = 2.2, Cx (600 km) = 2.5, Cx (800 km) = 

2.7, Cx (1000 km) = 2.75). 

The drag effect on the orbital elements is mainly a secular decrease on the semi-major axis: 




da 

dt (m/day) = -1.725.1012 a C S 

ρ X 

m 

with a (m), ρ (kg/m3 ), CX (m2 /kg). 

Figure 2.5-18 gives the the solar activity curve for the [01/2003-01/2015] period, taken into account in the da/dt 

and ∆V calculation (table 2.5-1) 

SOLAR ACTIVITY 01/2003-01/2015 

Figure 2.5-18: Solar activity for 01/2003-01/2015 period 

Table 2.5-1 gives the annual budget of the estimated altitude drift da/dt in m/year and the corresponding velocity 

increment ∆V (m/s) needed to compensate for this drift in order to maintain the nominal orbit for a 400-2000 km 

altitude range and for the years 2003 to 2014. The S/m coefficient depends on the size and mass of the payload, 

the local time of the ascending node, the orbit inclination, and the season. For the da/dt and DV calculation, the 

main hypothesis are given here after :The S/m ratio is equal to 10-2 m2 /kg 

the satellite mass is of 500 kg, 

the Cx coefficient is between 2.2 and 2.7, depending on the altitude, 

Year 




the orbit is Sun Synchronous with an ascending node at 12h00. 

the orbit maintenance strategy is based on semi-major axis correction (section 2.5.7.4). 

This table is given for information only; the calculation specific to the mission (taking into account the drag effective 

surface, the satellite mass, the chosen orbit, the orbit maintenance strategy) will be performed case by case. 

400 km 

450 km 

500 km 

550 km 

600 km 

650 km 

700 km 

750 km 

800 km 

900 km 

1000 km 

1100 km 

1200 km 

1500 km 

2000 km 

2003 2004 2005 2006 2007 2008 2009 2010 2011 2012 2013 2014 

198963 135517 77340 47854 44583 93321 240417 449580 525718 438443 278053 

112,5 76,7 43,8 27,1 25,2 52,8 136 254,3 297,4 248 157,3 82,3 

89583 57982 31422 17892 16658 37731 110102 221187 263154 216272 130138 

50,1 32,4 17,6 10 9,3 21,1 61,6 123,8 147,2 121 72,8 35,5 

42718 26292 13228 7138 6632 16552 53326 114078 139178 111964 64222 

23,6 14,6 7,3 4 3,7 9,2 29,5 63,1 77 62 35,5 15,9 

21355 12386 6028 3142 2868 7618 27000 61562 76158 60010 33320 14 

11,7 6,8 3,3 1,7 1,6 4,2 14,8 33,7 41,7 32,9 18,2 7,7 

12224 7239 3139 1551 1385 3878 15936 39222 48602 38114 19851 8 

6,6 3,9 1,7 0,8 0,8 2,1 8,6 21,2 26,3 20,6 10,8 4,3 

6365 3602 1568 784 747 2053 8399 21950 27773 21819 10919 39 

3,4 1,9 0,8 0,4 0,4 1,1 4,5 11,8 14,9 11,7 5,8 2,1 

3641 1886 830 509 490 1170 4565 12639 16619 12733 6169 216 

1,9 1 0,4 0,3 0,3 0,6 2,4 6,7 8,8 6,8 3,3 1,1 

1945 1049 553 343 324 591 2631 7550 9952 7492 3451 1182 

1 0,6 0,3 0,2 0,2 0,3 1,4 4 5,2 3,9 1,8 0,6 

1272 751 385 231 250 385 1638 4970 6704 4855 2177 848 

0,7 0,4 0,2 0,1 0,1 0,2 0,8 2,6 3,5 2,5 1,1 0,4 

531 374 216 138 138 236 747 2105 2616 1908 885 393 

0,3 0,2 0,1 0,1 0,1 0,1 0,4 1,1 1,3 1 0,5 0,2 

261 221 141 100 100 120 341 944 1285 964 442 221 

0,1 0,1 0,1 0,1 0,1 0,1 0,2 0,5 0,6 0,5 0,2 0,1 

164 143 61 61 61 102 225 451 615 512 266 143 

0,1 0,1 < 0,1 < 0,1 < 0,1 0,1 0,1 0,2 0,3 0,2 0,1 0,1 

125 63 63 42 42 63 125 272 397 272 167 63 

0,1


2.5.6.4 Solar radiation pressure 

The solar wind implies an alteration of the satellite momentum. 

The acceleration corresponding to the direct pressure is: 

S 

M = −K. 

. i . 

m 

p usun 

where m is the spacecraft mass, S the equivalent surface (spacecraft + solar arrays) perpendicular to the solar flux, K 

the radiation coefficient and equal to 4.56 10 -6 .N/m 2 , i p the illumination parameter (1 if the spacecraft is illuminated, 

0 if not), u sun the unit vector spacecraft-Sun. 

The main parameter is the angle β between the orbital plane and the Earth-Sun direction; for a given altitude, it 

determines the illuminated part x of the orbit (associated period Tβ). 

The usual effects on the orbital elements are long term perturbations on the eccentricity vector coordinates ex and ey, on the inclination i, on the position on orbit α. 

Typical values of this perturbation for the main orbital elements are: 

di 

dt = -5.2 10-4 degrees/year ; 

de 

dt 

y 

= 10 -4 /year 

d p so 

-3 = 3.10 degrees/year 

dt 

For Sun synchronous orbits, these terms become secular terms depending on the local hour of the ascending node 

and on x. 

2.5.6.5 Synthesis 

In comparison with the central term of Earth potential, the Earth oblateness term (J2) is the most important orbit 

perturbation. The main effects are secular drifts of the orbital parameters ω, Ω (typically a few degrees/day) and α. 

For long range studies, the non periodic variations of elements (secular terms) are predominant. 

The atmospheric drag causes a secular drift on the semi major axis which becomes very important under 600 km. 

For most missions, other effects can be considered as negligible in a first approach. 




2.5.7 ORBITAL MANOEUVRES 

The satellite is able to modify the orbit characteristics with its propulsion system. It can perform orbital manoeuvres 

for several reasons: 

in order to correct certain orbital parameters after injection by the launch vehicle (possible errors at injection 

or sometimes because the launch vehicle cannot deliver directly the satellite on its operational orbit), 

because of perturbations which are the result of a non ideal keplerian movement (see paragraph 2.5.4), 

for orbit transfer or rendez-vous manoeuvres. 

2.5.7.1 PROTEUS capabilities 

The PROTEUS Satellite manoeuvres use a propelling system with hydrazine (maximum capacity of the tank: 28 kg). 

In order to estimate the mass of propellant used for an instantaneous thrust necessary for orbital manoeuvres, the 

following formula is used: 

0 1 (1) 

 

−∆V 

 

g Isp 

∆m 

= m − e 

 

 

with mo is the initial mass of the satellite, Isp is the specific impulse of the engines (in seconds). 

Figure 2.5-19 shows the satellite capability over its life time in term of total velocity increment that allows a hydrazine 

mass of 28 kg. The curve gives the evolution of the ∆V versus payload mass. The assumptions for this chart are the 

following: 

a specific impulse I sp=220 s, 

a satellite initial mass m 0 = 330 + m ePL, assuming that m ePL is the equipped payload mass and that 330 

includes 28 kg of Hydrazine (see Table 3.1-1). 




Velocity Increment [m/s] 

190 

170 

150 

130 

110 

90 

0 50 100 150 

Equipped Payload Mass [kg] 

200 250 300 

Figure 2.5-19: Maximum DV as a function of the equipped payload mass 

A specific and detailed fuel budget will be calculated for each mission taking into account launch vehicle dispersions, 

the manoeuvres needs (established by mission analysis) and the performance criteria of the propulsion and its 

components (from Design and Acceptance testing). 




2.5.7.2 Cost in ∆V to change orbital parameters 

Orbital manoeuvres mainly consist in correcting the orbital parameters like the inclination and the semi major axis. 

This allows for control of the orbital plane drift, satellite position and phasing parameters. 

The cost in ∆V corresponding to an inclination or semi major axis correction in the case of low circular orbits can be 

evaluated using the charts provided hereafter. 

The orbital parameters variations (∆a, ∆e x, ∆e y, ∆i, ∆Ω, ∆α) can be expressed as functions of the tangential, radial 

and normal coordinates of the velocity increment ∆V: 

∆ 

∆a 

a VT 

= 2 (2) 

V 

∆VN 

∆i 

= cos α 

V 

∆VT ∆ex 

= 2cosα V 

∆VR 

+ sinα 

V 

∆ 

∆Ω = sinα 

VN 

sin i V 

∆VT ∆VR 

∆ey 

= 2sinα −cosα 

V V 

∆ 

∆α 

=−2 α ∆ 

− 

V 

V 

sin V 

tan i V 

R N 

Figure 2.5-20 gives the velocity increment for altitudes between 400 and 1600 km needed for semi major axis 

corrections (the formula (2) is applied for the calculation). For instance, if the need is to correct the semi major axis of 

40 km at an altitude equal to 700 km, this manoeuvre corresponds to a change in velocity ∆V = 21 m/s. It can be 

deduced from Figure 2.5-20 that the PROTEUS satellite can perform this manoeuvre with a payload mass up to 400 

kg. The associated consumed propellant mass can be estimated with formula (1). 




Figure 2.5-20: Semi major axis correction cost 

Figure 2.5-21 gives the velocity increment for altitudes between 400 and 1600 km, for an inclination correction from 

0 to 0.9 deg. Using the same process as above, one can estimate PROTEUS capacity in terms of inclination 

correction. 

Figure 2.5-21:Inclination correction cost 




2.5.7.3 Orbit positioning 

In most of the cases, after launch, a certain period is dedicated to the satellite in order for it to reach the target orbit 

necessary for mission completion. The initial orbit, depending on the capacity and on the accuracy of the launch 

vehicle, can be close to the operational orbit, or the satellite can be put on a parking orbit which can be quite 

different from the one of the mission itself. In both cases, the satellite must perform manoeuvres to be positioned on 

its nominal orbit. These positioning manoeuvres can modify every orbital parameter. In order to limit the propellant 

consumption of the satellite, the ∆V manoeuvres are performed tangential to the velocity (semi major axis 

corrections). 

Launch errors 

The satellite can not reach the target orbit with the sole launch vehicle; errors occur during flight and at the 

satellite/launch vehicle separation. These errors are usually estimated by a covariance matrix at injection. In order to 

correct the orbit parameters, two approaches can be applied: 

the satellite is close to the target orbit and the significant launch vehicle errors are corrected, 

the satellite is on a drift or parking orbit and the strategy consists in correcting the injection errors during the 

optimisation of the global orbit positioning process to decrease the manoeuvres number and the cost in ∆V. 

Strategy 

The strategy for orbit positioning often foresees a rendez-vous to achieve the target orbit with accuracy. In a first step, 

the strategy consists in introducing a sequence of impulse manoeuvres and Hohmann transfers between intermediate 

orbits in order that the satellite reaches the target orbit. 

As soon as the User chooses the sequence, the kind and the number of orbital manoeuvres, he can estimate the cost 

in ∆V, and so the propellant consumption of the satellite and the mission feasibility with a PROTEUS system can then 

be deduced (see paragraphs 2.5.6.1 and 2.5.6.2). 

2.5.7.4 Orbit maintenance 

For most of the missions a station keeping strategy is necessary to take into account every constraint. For instance, 

the phased orbits need to have a well defined semi major axis. In this case, because of orbital perturbation effects, 

the semi major axis value must be regularly corrected to maintain the satellite in a given altitude range. The allowed 

variations of orbital parameters depend on the mission and their limits can be deduced from a specific analysis. 

The usual strategy consists in correcting the semi major axis. The inclination parameter is less often altered. The 

number and the amplitude of manoeuvres depend on the strategy applied to position the satellite in a given window. 

The station keeping manoeuvres can affect every orbital parameters (a, e x, e y, i, Ω, α). The main perturbation 

concerning orbit maintenance is due to the atmospheric drag. As soon as the altitude and the duration of the mission 

are known, one can estimate the cost in ∆V needed to compensate for atmospheric drag (see Table 2.5-1) and the 

corresponding inclination correction can be evaluated with Figure 2.5-1. 




2.5.7.5 Synthesis 

The main orbital manoeuvres during orbit positioning and maintenance are used to compensate for launch errors 

dispersion, phasing if necessary, and atmospheric drag. The most usual manoeuvres consist in correcting the semi 

major axis and the inclination. Each mission needs a different strategy depending on the mission constraints, the 

satellite AOCS properties, and the station visibility. For instance, a mission can require to perform every orbital 

manoeuvre in eclipse or over the sea while the payload is turned off. These constraints, once included into the 

optimisation process, can modify the manoeuvre schedule. The manoeuvre frequency mainly depends on the satellite 

altitude and on the window allowed for the orbital parameters. 

2.5.8 ORBIT DEBRIS GENERATION ANALYSIS 

Analysis shall be led and documented for each Proteus mission to assess orbital debris generation potential and 

debris mitigation options. 

This analysis is required in particular to demonstrate compliance with the requirements of NASA Directive NPD 

8710.3 and [RD12] 

The analysis shall include the following: 

- potential for orbital debris generation in both nominal operation and malfunction conditions, including 

malfunctions during launch phase. 

- potential for orbital debris generation due to on-orbit impact with existing space debris (natural or human 

generated) or other orbiting space systems. 

Such orbital debris generation analysis was performed for the JASON 1 mission [RD13], and can be used as 

reference for subsequent analysis reports. 

The Payload Supplier shall provide input for the corresponding Satellite System Analysis. 





Chapter 3 : Payload interface requirements 



N°§ PUID Change 

Status 

§3 Modified 

in 

§3 Modified 

in 

Reason of Change Change Reference Doc 

Issue 

STR cable introduction CIIS.4.1.JC.1_1 6.2 

STR cable introduction CIIS.4.1.JC.1_1 6.2 

§3.1 New in 6.2 

§3.1.1.1 [PL - 3.1.1 - 1 a] Modified 

in 

§3.1.1.1 Modified 

in 


in 

Equipped paylaod mass updated PUM.6.1.CG.31_20 6.2 

Platform mass updated, Launch 

vehicle adapter mass updated 



CIIS.4.1.JC.1_2 6.2 

Reference to Figure 3.1-3 removed CIIS.4.1.JC.1_2 6.2 

§3.1.1.1 New in STA position on the payload CIIS.4.1.JC.1_2 6.2 


in 

§3.1.1.2 [PL - 3.1.1 - 3 a] Modified 

in 

§3.1.1.2 [PL - 3.1.1 - 4 a] Modified 

in 


in 


in 


in 

§3.1.4.1.2.1 [PL - 3.1.4 - 3 a] Modified 

in 

§3.1.4.1.2.2 [PL - 3.1.4 - 4 a] Modified 

in 

"mission" replaced by "launch vehicle" CIIS.4.1.JC.1_2 6.2 

Reference to Figure 3.1-3 removed CIIS.4.1.JC.1_2 6.2 

CoG location accuracy on complete 

FM 

PUM.6.2.EJ.03 6.2 

Reference to Figure 3.1-3 removed CIIS.4.1.JC.1-2 6.2 

Clarification on counterbalance 

masses considered 

CIIS.4.1.JC.1-2 6.2 

Figure 3.1-2 title updated CIIS.4.1.JC.1-2 6.2 

Flatness requirement updated PUM.6.1.CG.31_1 6.2 

Parallelism requirement updated PUM.6.1.CG.31_2 6.2



Status 


Issue 

§3.1.4.2.1 Figure 3.1-13 updated PUM.6.1.CG.31_3 6.2 

§3.1.4.2.1 New in Nota added PUM.6.1.CG.31_3 6.2 

§3.1.4.2.1 New in H02, H03 bracket masses PUM.6.1.CG.31_20 6.2 

§3.1.4.3.2.2 Modified 

in 

Title modified PUM.6.1.CG.31_4 6.2 

§3.1.4.3.2.2 New in Reference to Figure 3.6-2b PUM.6.1.CG.31_4 6.2 

§3.2.2.2 [PL - 3.2.2 -8 ] New in Constraints on (C1,C2) thermal 

parameters 

§3.4.1 New in Correspondance between P1, P2 & 

P3 with SiOP 3, SiOP 4, SiOP 5 



6.2 

CIIS.4.1.JC.2_1 6.2 

§3.4.3.1 New in P/F test point availability PUM.6.1.CG.31_5 6.2 

§3.4.3.1 [PL - 3.4.3 - 6 a] Modified 

in 

Up to 16 (instead of 12) units to be 

connected 

§3.4.5.1 New in Rationale of the PL-3.4.5-1 

requirement 

§3.4.5.3.1.2 Modified 

in 


in 

§3.4.6.2 [PL - 3.4.6 - 6 ] Deleted 

in 

§3.4.6.3.1 [PL - 3.4.6 – 7 a] Modified 

in 

§3.4.6.3.1 Deleted 

in 

§3.4.7.1 [PL - 3.4.7 - 2 a] Modified 

in 

§3.4.7.1 [PL - 3.4.7 - 4 a] Modified 

in 


in 


in 

§3.4.7.1 [PL - 3.4.7 - 3 a] Modified 

in 

PUM.61.CG.31_13 6.2 

PUM.6.1.OR.2 6.2 

Nota displaced PUM.6.1.CG.31_9 6.2 

Lines "8" and "16" which may be ON 

precised 

P/F actions definition in case of PL 

anomaly removed 

PUM.6.1.CG.31_10 6.2 

PUM.6.2.CG.31_10 6.2 

CIIS.4.1.JC.4_5 6.2 

Figure 3.4-7 removed CIIS.4.1.JC.4_5 6.2 

New wording PUM.6.1.CG.31_11 6.2 


Bit allocation for time distribution 

updated 

Bit allocation for time distribution 

updated 

PUM.6.1.CG.31_11 6.2 

PUM.6.1.CG.31_11 6.2 


§3.4.7.1 New in Figure 3.4-4 added PUM.6.1.CG.31_11 6.2 

§3.5.2 New in Connector designation (fixed or 

mobile) 

CIIS.4.2.JC.1_5 6.2 

§3.5.2.2 Change of connector reference 6.2



Status 


in 

§3.5.3.1 [PL - 3.5.3 - 1 a] Modified 

in 


in 


Issue 

Change of connector reference PUM.6.1.EJ.31 6.2 

Reference to voltage removed CIIS.4.1.JC.1_9 6.2 

Reference in Figure 3.5-7 to voltage 

removed 

§3.5.3.1 New in Correspondance P1,P2 & P3 with 

SiOP 3, SiOP 4 and SiOP 5 

§3.5.3.2 [PL - 3.5.3 - 3 a] Modified 

in 

§3.5.3.2 [PL - 3.5.3 - 4 a] Modified 

in 

§3.5.3.3.1 Modified 

in 

§3.5.3.3.1 [PL - 3.5.3 - 6 a] Modified 

in 

§3.5.3.3.3 [PL - 3.5.3 - 8 ] Modified 

in 


in 

§3.5.3.3.5 [PL - 3.5.3 - 11 

a] 

Modified 

in 


in 

§3.5.3.3.8 [PL - 3.5.3 - 16 

a] 

Modified 

in 


in 


in 


in 


in 

Reference to nominal voltage 

removed 



CIIS.4.1.JC.1_9 6.2 

CIIS.4.1.JC.2_1 6.2 

CIIS.4.1.JC.1_16 6.2 

28 V voltage added CIIS.4.1.JC.1_16 6.2 

Nominal voltage removed CIIS.4.1.JC.1_16 6.2 

Nominal and degraded voltage 

ranges defined 

CIIS.4.1.JC.1_16 6.2 

Text re-introduced in the requirement CIIS.4.1.JC.1_17 6.2 

One sentence removed CIIS.4.1.JC.1_17 6.2 

Compatibility with the fuse blowing CIIS.4.1.JC.1_17 6.2 

Figure 3.5-29 added CIIS.4.1.JC.1_17 6.2 

Minimim voltage defined in PL-3.5.3- 

6 


6 


6 

Coherence between Note 1 and PL- 

3.5.9-2 

Coherence between Note and PL- 

3.5.9-2 

CIIS.4.1.JC.1_16 6.2 

CIIS.4.1.JC.1_16 6.2 

CIIS.4.1.JC.1_16 6.2 

CIIS.4.1.JC.1_6 6.2 

CIIS.4.1.JC.1_6 6.2 

§3.6.2.1 Figure 3.6-2 Title corrected PUM.6.1.CG.31_13 6.2 


in 


in 

Figure 3.6-2b updated PUM.6.1.CG.31_13 6.2 

STA optical cube positioned in Figure 

3.6-3 

PUM.6.1.CG.31_16 6.2 

§3.6.2.2.5 New in SRA definition mission dependant PUM.6.1.CG.31_15 

a 

6.2



Status 


in 


Issue 

STA stiffness specified PUM.6.1.CG.31_15 

a 

§3.6.2.2.5 New in STA FEM model to be provided to the 

PL 

§3.6.2.2.8 [PL - 3.6.2 - 8 a] Modified 

in 

PUM.6.1.CG.31_15 

a 



6.2 

6.2 

STR clearance angles updated CIIS.4.1.JC.1_12 6.2 

§3.6.2.2.8 New in Moon effect to be analysed CIIS.4.1.JC.1_12 6.2 

§3.6.2.3.2 [PL - 3.6.2 - 12 

a] 

Modified 

in 

STA random vibration levels updated PUM.6.1.CG.31_17 6.2 

§3.6.2.3.2 Table 3.6-2 updated PUM.6.1.CG.31_17 6.2 

§3.6.2.3.2 New in Figure 3.6-6 added PUM.6.1.CG.31_17 6.2 

§3.6.2.3.2 New in Additional sentence PUM.6.1.CG.31_17 6.2 

§3.6.3 Modified 

in 


in 


in 

§3.6.3 [PL - 3.6.3 - 2 a] Modified 

in 

§3.6.3 [PL - 3.6.3 - 3 a] Modified 

in 

Introduction of H20 connector 

bracket 

Figure 3.6-5 updated (introduction of 

H20 connector bracket) 

PUM.6.1.CG.31_18 6.2 

PUM.6.1.CG.31_18 6.2 

Suppression of reference to line N°14 PUM.6.1.CG.31_18 6.2 

Wiring routing of the thermal control 

harness 

Introduction of Anchor #1 & 2 on 

STA baseplate 

§3.6.3 New in Introduction of Anchor #1 & 2 on 

STA baseplate 


in 

Figure 3.6-7 added (Introduction of 

Anchor #1 & 2 on STA baseplate) 

§3.6.3 New in Table 3.6-4 added (STR cable 

lengths) 


in 

§3.6.4 [PL - 3.6.4 - 1 a] Modified 

in 

PUM.6.1.CG.31_19 6.2 

PUM.6.1.CG.31_20 6.2 

PUM.6.1.CG.31_20 6.2 

PUM.6.1.CG.31_20 6.2 

PUM.6.1.CG.31_20 6.2 

STR wires masses PUM.6.1.CG.31_20 6.2 

updated tension in the 8 interface 

screws 

RID 

CIIS.4.1.JC.1_13 

§3.6.4 [PL - 3.6.4 -2 ] New in Interface screws type RID 

CIIS.4.1.JC.1_13 

§3.6.4 [PL - 3.6.4 -3 ] New in Tightening torque of the interface 

screws to be defined and justified 

§3.6.4 [PL - 3.6.4 -4 ] New in Payload insert mechanical loads to 

support the STA 

§3.6.4 New in Data to be provided by the Payload 

Supplier 

RID 

CIIS.4.1.JC.1_13 

RID 

CIIS.4.1.JC.1_13 

RID 

CIIS.4.1.JC.1_13 

6.2 

6.2 

6.2 

6.2 

6.2



Status 

§3.7 Modified 

in 

§3.7.1.2.3.3 [PL - 3.7.1 - 13 

a] 

§3.7.1.2.4 [PL - 3.7.1 - 15 

a] 

§3.7.2.2.4 [PL - 3.7.2 - 18 

a] 

Modified 

in 

Modified 

in 

Modified 

in 


Issue 

Additional sentence PUM.6.1.CG.31_21 6.2 

Precision on static tests and 

inspection report added 



Status 



PUM.6.1.CG.31_21 6.2 

Certificate of conformity precised PUM.6.1.CG.31_21 6.2 

Certificate of conformity precised PUM.6.1.CG.31_22 6.2 

Reason of Change Change 

Reference 

§3.1.5.1 [PL - 3.1.5 - 0 ] New in Clarification PUM.6.2.EJ.04 6.3 

§3.2.2.2 [PL - 3.2.2 -8 

a] 

Modified 

in 


in 

Constraints on (C1,C2) thermal 

parameters 

Doc 

Issue 

PUM.6.1.EJ.34a 6.3 

Coherence with generic PAYLINT packet PUM.6.2.EJ.25 + 

PRTS-DIM-0043 

§3.4.2 Coherence with generic PAYLINT packet PUM.6.2.EJ.25 + 

PRTS-DIM-0043 

§3.4.3 [PL - 3.4.3 -1 

a] 

Modified 

in 

6.3 

6.3 

1 instrument is a 1553 RT PUM.6.2.EJ.28a 6.3 

§3.4.3.4 [PL - 3.4.3 -21 ] New in Clarification PUM.6.2.EJ.05 6.3 

§3.4.4.3.1.2 [PL - 3.4.4 - 6 

a] 

Modified 

in 

Subaddress 16d added RID N° 

CIIS4.1.JC1_4a 

§3.4.4.3.1.4 New in Average data rate is mission dependent PUM.6.2.TH.02 6.3 

§3.4.5.3.1.1 New in Risk of anomaly on PLYTM due to remain 

on elephant packets 

§3.4.5.3.1.1 Modified 

in 

§3.4.6 [PL - 3.4.6 - 1 

a] 

Modified 

in 

6.3 

PUM.6.2.PL.01 6.3 

Clarification RID N° 

CIIS4.1.JC1_4a 

6.3 

PL surveillance clarification PUM.6.2.EJ.28a 6.3 

§3.4.6 New in PL surveillance clarification PUM.6.2.EJ.28a 6.3 


in 


§3.4.6 New in PL surveillance clarification PUM.6.2.EJ.28a 6.3 


in 


§3.4.6.1 New in PL surveillance clarification PUM.6.2.EJ.28a 6.3 

§3.4.6.2.1 New in New title PUM.6.2.EJ.28a 6.3



Status 


Reference 

§3.4.6.2.2 New in New title PUM.6.2.EJ.28a 6.3 

§3.4.6.2.2 [PL - 3.4.6 -6 

]a 

Modified 

in 



Doc 

Issue 


§3.4.6.2.2 New in PL surveillance clarification PUM.6.2.EJ.28a 6.3 

§3.4.6.3 [PL - 3.4.6 -7 ] 

b 

Modified 

in 


§3.4.6.3 [PL - 3.4.6 -9 ] New in Tuning rules to establish the thresholds 

for each surveillance 

PUM.6.2.PL.02 6.3 

§3.4.6.3 New in Figure 3.4-7 added PUM.6.2.EJ.28a 6.3 

§3.5.2 New in Lines not used by PL PUM.6.2.EJ.31 6.3 

§3.5.3.1 [PL - 3.5.3 –1b 

] 

Modified 

in 

one instrument correponds to one power 

line 

§3.5.3.1 New in Configuration not recommended during 

launch 

PUM.6.2.EJ.28a 6.3 

CIIS.4.1.JC.1_9a 6.3 

§3.5.4.1 New in Safe plug arm PUM.6.2.EJ.27 6.3 

§3.5.4.1 New in Pyro lines activition PUM.6.2.EJ.27 6.3 

§3.5.4.1 [PL - 3.5.4 -12 ] New in Additional pyro lines requirement PUM.6.2.EJ.27 6.3 





§3.5.4.2 New in PUM.6.2.EJ.27 6.3 










§3.5.6 [PL - 3.5.6 -28 ] New in PL OFF to be compaible with DHU 

outputs 

PUM.6.2.EJ.25 + 

PRTS-DIM-0043 

§3.5.6.1.2 [PL - 3.5.6 -18 ] New in To be documented in IDS/ICD PUM.6.2.EJ.06 6.3 


in 

Complinace with DHU IDS PUM.6.2.EJ.23 6.3 

6.3



Status 


in 


Reference 



Doc 

Issue 

Compliance with DHU IDS PUM.6.2.EJ.23 6.3 



in 


in 



§3.5.6.2.1 [PL - 3.5.6 -20 ] New in Compatibility with DHU active analog 

input 


in 


in 

PUM.6.2.EJ.06 6.3 







in 




in 


in 





§3.5.6.4 [PL - 3.5.6 -27 ] New in To be documented in IDS/ICD PUM.6.2.EJ.06 6.3 


in 

DELTA II radiated field levels updated PUM.6.2.EJ.29 6.3 

§3.5.7.2.2 New in DELTA RF environment on launch site 

defined 

PUM.6.2.EJ.29 6.3 

§3.5.7.2.2 New in ROCKOT radiated field updated PUM.6.2.EJ.29 6.3 

§3.5.7.2.2 New in Figure which displays the RF environment 

added 

§3.5.7.2.2 New in SOYUZ belongs to the selected launch 

vehicles 


vehicles 


vehicles 

PUM.6.2.EJ.29 6.3 

PUM.6.2.EJ.29 6.3 

PUM.6.2.EJ.29 6.3 

PUM.6.2.EJ.29 6.3



Status 


in 


Reference 



Doc 

Issue 

Standard STA IDS PUM.6.2.TH.01 6.3 

§3.6.1 New in Standard STA compatibility with PL 

interface in carbon 

§3.6.1 New in Compatibility of the standard STA to be 

assessed 

PUM.6.2.TH.01 6.3 

PUM.6.2.TH.01 6.3 

§3.6.1 New in STA adaptation PUM.6.2.TH.01 6.3 

§3.6.2.3.1 [PL - 3.6.2 - 9 

a] 

Modified 

in 


in 


in 

To be defined at STA I/F level PUM.6.2.EJ.32 6.3 



§3.6.2.3.2 New in To be defined at STA I/F level PUM.6.2.EJ.32 6.3 

§3.6.6 New in STA ground braids to be provided by PF CIIS.4.1.JC.3_1 6.3 

§3.6.6 [PL - 3.6.6 -1 ] New in New Section: STA grounding on PL CIIS.4.1.JC.3_1 6.3 

§3.7.1.2.2.3 Modified 

in 

Factors of safety modified PUM.6.2.EJ.24 6.3



3. PAYLOAD INTERFACE REQUIREMENTS 17 

3.1 MECHANICAL INTERFACE REQUIREMENTS 17 

3.1.1 MASS PROPERTIES 18 

3.1.1.1 Mass 18 

3.1.1.2 Centering 18 

3.1.1.3 Inertia 21 

3.1.2 STIFFNESS 22 

3.1.2.1 Stiffness in launch configuration 22 

3.1.2.2 Stiffness in flight configuration 23 

3.1.3 AVAILABLE VOLUME FOR THE PAYLOAD 24 

3.1.3.1 Launch vehicles fairing constraint 24 

3.1.3.2 In flight configuration constraint 25 

3.1.4 MECHANICAL INTERFACES PAYLOAD/PLATFORM 27 

3.1.4.1 External interfaces for the payload accommodation 27 

3.1.4.2 Connector brackets interfaces 32 

3.1.4.3 Star Trackers Assembly interfaces 38 

3.1.5 MAXIMUM GENERATED DISTURBANCES 39 

3.1.5.1 Dynamic disturbances 39 

3.1.5.2 Maximum Shock generated by the payload 39 

3.1.6 OPTIONAL PAYLOAD MODULE 39 

3.2 THERMAL INTERFACE REQUIREMENTS 41 

3.2.1 PLATFORM-PAYLOAD CONDUCTIVE AND RADIATIVE INTERFACES 42 

3.2.2 ACTIVE THERMAL CONTROL 45 

3.2.2.1 Heaters lines and thermistors 45 

3.2.2.2 Regulation algorithm 47 

3.2.3 PAYLOAD THERMAL MONITORING 49 

3.3 POWER SUPPLY INTERFACE REQUIREMENTS 50 

3.3.1 MEAN ORBITAL POWER AVAILABLE FOR THE PAYLOAD 50 

3.3.2 POWER PEAKS LIMITATIONS FOR THE PAYLOAD 52 

3.3.3 POWER SUPPLY DURING TRANSIENTS 53 

3.3.3.1 Launch phase 53 

3.3.3.2 SHM phase 53 

3.3.3.3 Orbit Control phase 53 

3.4 COMMAND & CONTROL INTERFACE REQUIREMENTS 54 

3.4.1 COMMAND AND CONTROL AVAILABLE LINES 54 

3.4.2 PAYLOAD INSTRUMENT COMMAND/CONTROL STATUS 56 

3.4.3 MIL-STD-1553B DATA BUS INTERFACE 57 

3.4.3.1 System requirements 58 

3.4.3.2 Description of payload units behaviour 60 

3.4.3.3 Protocol general requirements 60 

3.4.3.4 Data Bus interface characteristics 61 

3.4.3.5 Initialisation of the protocol 63 

3.4.4 PAYLOAD COMMANDABILITY 64 

3.4.4.1 General 64 

3.4.4.2 Discrete commands 64 

3.4.4.3 1553 commands 65 

3.4.5 PAYLOAD TELEMETRY 68 

3.4.5.1 General 68 

3.4.5.2 TM from discrete acquisitions 68 




3.4.5.3 TM from 1553 acquisitions 69 

3.4.5.4 Deleted 74 

3.4.6 PAYLOAD SURVEILLANCE 75 

3.4.6.1 Failure Detection Isolation and Recovery 75 

3.4.6.2 Payload switch-off in case of SHM 76 

3.4.6.3 Payload switch-off in case of payload anomaly 78 

3.4.7 TIME AND SYNCHRONISATION DISTRIBUTION ERREUR! SIGNET NON DÉFINI.79 

3.4.8 PAYLOAD SPECIFIC SOFTWARE INSIDE DHU 81 

3.5 ELECTRICAL INTERFACE REQUIREMENTS 82 

3.5.1 GENERAL SYSTEM CONFIGURATION 82 

3.5.2 PIN ALLOCATION 84 

3.5.2.1 Power bracket 84 

3.5.2.2 Acquisition and command interface brackets 85 

3.5.3 POWER LINES 87 

3.5.3.1 Available lines 87 

3.5.3.2 Payload power consumption 88 

3.5.3.3 Power interface characteristics 88 

3.5.4 PYROTECHNIC LINES 93 

3.5.5 THERMAL LINES 96 

3.5.5.1 Active thermal control 96 

3.5.5.2 Thermal monitoring 96 

3.5.6 COMMAND AND CONTROL LINES 97 

3.5.6.1 Commands 97 

3.5.6.2 Telemetry 106 

3.5.6.3 Time distribution and synchronization 114 

3.5.6.4 MIL-STD-1553B bus 115 

3.5.6.5 Deleted 115 

3.5.7 ELECTROMAGNETIC INTERFACE REQUIREMENTS 116 

3.5.7.1 Conducted Emission & Susceptibility Requirements 116 

3.5.7.2 Radiated Emission and Susceptibility Requirements 123 

3.5.8 ESD PROTECTION 129 

3.5.8.1 Direct arc discharge 129 

3.5.8.2 Indirect arc discharge 129 

3.5.9 MAGNETIC FIELD INTERFACE REQUIREMENTS 130 

3.5.9.1 Emission requirements 130 

3.5.9.2 Susceptibility requirements 130 

3.6 STAR TRACKER ASSEMBLY ACCOMMODATION 132 

3.6.1 GENERAL 132 

3.6.2 MECHANICAL SPECIFICATIONS 134 

3.6.2.1 Interfaces 134 

3.6.2.2 Physical characteristics 136 

3.6.2.3 Dynamic Environment 141 

3.6.3 HARNESS CONSTRAINTS 145 

3.6.4 THERMAL DESIGN AND INTERFACE REQUIREMENTS 148 

3.6.5 CLEANLINESS REQUIREMENTS 149 

1493.6.6 STA GROUNDING ON PAYLOAD 149 

3.7 GROUND SUPPORT EQUIPMENT INTERFACES 151 

3.7.1 MECHANICAL GSE INTERFACES 151 

3.7.1.1 General 151 

3.7.1.2 Requirements for delivered MGSE 151 

3.7.2 ELECTRICAL GSE INTERFACES 155 

3.7.2.1 General 155 

3.7.2.2 Requirements for delivered EGSE 155 





Figure 3.1-1 : Satellite CoG Xs location for different equipped payload masses and for different equipped payload 

CoG Xp locations.......................................................................................................................................... 19 

Figure 3.1-2 : Equipped payload CoG allowed Yp/Zp locations in function of the equipped payload mass ............ 20 

Figure 3.1-4 : Stiffness requirements at equipped payload level............................................................................. 22 

Figure 3.1-5 : Payload volume under fairings........................................................................................................ 24 

Figure 3.1-6 : Payload allowed volume in flight configuration ............................................................................... 25 

Figure 3.1-7 : Volume occupied by the solar array rotation ................................................................................... 26 

Figure 3.1-8 : Platform Y view .............................................................................................................................. 27 

Figure 3.1-19 : Platform Z view ............................................................................................................................ 27 

Figure 3.1-9 : +Xs panel platform with its four PF/PL mechanical interfaces........................................................... 28 

Figure 3.1-10: Payload fitting interface ................................................................................................................. 29 

Figure 3.1-11 : Platform / payload interface details .............................................................................................. 29 

Figure 3.1-12 : Payload mounting interface detail, case of payload with the same structure as platform................. 30 

Figure 3.1-13 : H01 connector bracket mechanical interface................................................................................. 32 

Figure 3.1-14 : Electrical interface brackets global view ........................................................................................ 33 

Figure 3.1-15 : Electrical brackets volume............................................................................................................. 34 

Figure 3.1-16 : H02 & H03 electrical brackets : mechanical interface plane .......................................................... 36 

Figure 3.1-17 : H02 & H03 electrical brackets : connectors interface..................................................................... 36 

Figure 3.1-18 : Payload JASON module as example............................................................................................. 40 

Figure 3.2-3 : MLI typical interface between platform and payload........................................................................ 43 

Figure 3.2-1 : Typical dimensions of Platform thermal radiators ............................................................................ 44 

Figure 3.2-2 : Heaters and thermistors redundancy configuration for thermal control ............................................ 46 

Figure 3.3-1 : Maximal satellite mean orbital power w.r.t the altitude and the inclination ....................................... 51 

Figure 3.3-2 : Maximal satellite mean orbital power w.r.t. the ascending node of the sun synchronous orbit and the 

altitude ......................................................................................................................................................... 51 

Figure 3.3-3 : Maximal satellite mean orbital power versus the altitude for Sun-synchronous orbits LHAN 18 h ..... 52 

Figure 3.3-4 : Allocation for payload power consumption during SHM phase .......................................................53 

Figure 3.4-1 : Payload instrument command/control status ................................................................................... 56 

Figure 3.4-4: Timing of typical 1553 exchanges between platform and payload (corresponding respectively to few 

and many data transferred on the bus).......................................................................................................... 57 

Figure 3.4-6: interface between Payload Unit connected to 1553 and H02/H03.................................................... 59 

Figure 3.4-5: 1553 status word ............................................................................................................................ 61 

Figure 3.4-2: TC communication .......................................................................................................................... 66 

Figure 3.4-3 : Scientific Telemetry Exchange.......................................................................................................... 73 

Figure 3.4-7: Time bulletin versus PPS signal......................................................................................................... 80 

Figure 3.5-1 : System configuration with redundant units and cross-strapping ....................................................... 82 

Figure 3.5-2 : System configuration with single unit internally redundant ............................................................... 83 

Figure 3.5-3 : System configuration with no redundant unit ................................................................................... 83 

Figure 3.5-4 : H01 Connector bracket .................................................................................................................. 84 



Figure 3.5-7 : DHU I/O channel layout................................................................................................................. 87 

Figure 3.5-8 : DHU output impedance.................................................................................................................. 89 

Figure 3.5-29: Transients of the power bus to fuse blowing ................................................................................... 91 

Figure 3.5-9 : Electrical inhibit implementation (only the main branch is illustrated) ............................................... 91 

Figure 3.5-10 : Signal wave shape for the HLC pulses........................................................................................... 99 

Figure 3.5-11 :Electrical interface for the SBDL.................................................................................................... 100 




Figure 3.5-11a: Receiver input impedance (differential)....................................................................................... 101 

Figure 3.5-12 : Signal wave shape for the LLC pulses ........................................................................................ 101 

Figure 3.5-13 : Electrical interface for the SBDL................................................................................................... 103 

Figure 3.5-14 : Serial Command timing (B0 is MSB) ........................................................................................... 104 

Figure 3.5-15 : Electrical interface for the SBDL................................................................................................... 111 

Figure 3.5-16 : Digital Serial Acquisition (8 bit) timing......................................................................................... 112 

Figure 3.5-17 : Digital Serial Acquisition (16 bit) timing (B0 is MSB)..................................................................... 113 

Figure 3.5-19 : Conducted emission over the power supply bus (Narrowband).................................................... 116 

Figure 3.5-20 : Inrush Current profile ................................................................................................................. 117 

Figure 3.5-21 : Off-switching transient................................................................................................................ 118 

Figure 3.5-22 : Susceptibility to sine conducted emissions ................................................................................... 119 

Figure 3.5-23 : Conducted susceptibility, transient wave shape ........................................................................... 121 

Figure 3.5-26: Common mode voltage............................................................................................................... 122 

Figure 3.5-24 : Radiated emission, E-field, Narrow band .................................................................................... 123 

Figure 3.5-25 : Radiated susceptibility, E-field ..................................................................................................... 125 

Figure 3.5-31: ROCKOT Launch Vehicle RF environment .................................................................................... 126 

Figure 3.5-32: LV and Launch base Emission Spectra (Soyuz ST Configuration) ................................................... 128 

Figure 3.5-27: Unit under direct arc discharge.................................................................................................... 129 

Figure 3.5-28: Unit under indirect arc discharge ................................................................................................. 129 

Figure 3.6-1 : Standard Star Trackers Assembly .................................................................................................. 132 

Figure 3.6-2 : Standard Star Trackers Assembly interface plane........................................................................... 134 

Figure 3.6-2b : Interface cross section................................................................................................................. 135 

Figure 3.6-3 : Standard Star Trackers Assembly volume ...................................................................................... 136 

Figure 3.6-3c : F1 view of the STR 1 with reference cube orientation.................................................................... 138 

Figure 3.6-3d : F2 view of the STR 2 with reference cube orientation ................................................................... 139 

Figure 3.6-4 : Azimuth and elevation definition : CALIPSO (111° and 45°) .......................................................... 140 

Figure 3.6-6: Maximum random vibration levels at Star Trackers Assembly level.................................................. 143 

Figure 3.6-5 : STAs wiring (STRs and CTA) .......................................................................................................... 145 

Figure 3.6-7 : Anchor# 1 and Anchor# 2 positions on the STA baseplate ........................................................... 147 

Figure 3.6-8: Ground stud configuration for STA grounding................................................................................ 150 





Table 3.1-1: Satellite module masses.................................................................................................................... 18 

Table 3.1-2 : Interface mechanical distortions ....................................................................................................... 31 

Table 3.1-3 : Maximum Shock levels generated at the PF/PL I/F plane by the Payload ........................................... 39 

Table 3.2-1 : Thermo-optic characteristics of the platform parts ............................................................................ 42 

Table 3.4-1: Distribution of the Command & Control resources inside the DHU.....................................................54 

Table 3.4-2: "mode_code" commands usable by the Payload ................................................................................ 62 

Table 3.4-3: Subaddresses of the address 31 reserved for broadcast command management ............................... 62 

Table 3.4-4: Payload switch-off sequence by reconfiguration module (see figure 3.5-7 for lines group definition)... 76 

Table 3.5-1 : H01 connector bracket description................................................................................................... 84 



Table 3.5-4 : Electrical characteristics of the DHU HLC output interface................................................................. 98 

Table 3.5-5 : USER High Level Command input interface ...................................................................................... 99 

Table 3.5-6 : Electrical characteristics of the SBDL complementary driver............................................................. 100 

Table 3.5-7 : Electrical characteristics of the SBDL complementary receiver.......................................................... 101 

Table 3.5-8a: DRIVER and RECEIVER values for Data, Clock and enable ............................................................. 103 

Table 3.5-8 :Electrical characteristics of the SBDL complementary driver.............................................................. 103 

Table 3.5-9 : Electrical characteristics of the SBDL complementary receiver.......................................................... 104 

Table 3.5-10 : Characteristic Times values.......................................................................................................... 105 

Table 3.5-11 : Electrical characteristics of the DHU AN input interface................................................................. 106 

Table 3.5-12 : Analog monitoring output interface (USER side)............................................................................ 106 

Table 3.5-13 : Electrical characteristics of the DHU TH input interface ................................................................. 108 

Table 3.5-14 : Interconnection characteristics ..................................................................................................... 108 

Table 3.5-15 : Output characteristics .................................................................................................................. 108 

Table 3.5-16 : Electrical characteristics of the DHU DR input interface ................................................................. 109 

Table 3.5-17 : Digital relay monitoring output interface (USER side) .................................................................... 109 

Table 3.5-17a : DRS protocol in S/W register...................................................................................................... 109 

Table 3.5-18 : Electrical characteristics of the DHU DB input interface ................................................................. 110 

Table 3.5-19 : Digital bilevel monitoring output interface (USER side) .................................................................. 110 

Table 3.5-20a: DRIVER and RECEIVER values for Data, Clock and enable ........................................................... 111 

Table 3.5-20 : Electrical characteristics of the SBDL complementary driver........................................................... 111 

Table 3.5-21 : Electrical characteristics of the SBDL complementary receiver........................................................ 112 

Table 3.5-22 : Digital Serial Acquisition timing.................................................................................................... 113 

Table 3.5-24 : Requirements about radiated emission, E field and narrow band.................................................. 124 

Table 3.5-28 ROCKOT L/V transmitters radiated field levels................................................................................ 126 

Table 3.5-29: LV and launch base mission Spectra (Soyuz ST configuration)........................................................ 127 

Table 3.5-25 Volume C1 where the magnetic field is between 1 and 3 Gauss in satellite nominal mode.............. 130 

Table 3.5-26 Volume C2 where the magnetic field can reach 23 Gauss in satellite SHM...................................... 130 

Table 3.6-1 : Quasi static acceleration loads ...................................................................................................... 141 

Table 3.6-2 : Maximum random vibration levels at Star Trackers Assembly level.................................................. 142 

Table 3.6-3 : Maximum Shock levels at Star Trackers Assembly level ................................................................... 144 

Table 3.6-4: STR cables length ........................................................................................................................... 148 

Table 3.7-1 : Factors of safety ............................................................................................................................ 153 







LIST OF FIGURES ................................................................................................................................................. 11 

LIST OF TABLES.................................................................................................................................................... 13 

LIST OF CHANGE TRACEABILITY .......................................................................................................................... 14 




LIST OF TBCs 


Resolution 

§3.1.1.3.1 In launch configuration, the absolute values of the equipped payload inertia 

expressed in the Payload Reference Frame (O P, X P, Y P, Z P) F p shall be less than 75 

kg.m 2 relative to the (O P, X P) axis and less than 300 kg.m 2 (TBC) relative to the two 

other axes (O P, Y P) and (O P, Z P). 

§3.1.2.2 The first mode frequencies allowed at equipped payload level in flight 

configuration shall be higher than 5 Hz (TBC). 

§3.1.4.1.2.1 The relative flatness between the four mounting surfaces provided by the platform 

will be better than 0.1 mm (obtained by manufacturing or wedging) (TBC). 

§3.3.3.2 Payload power consumption (including thermal control) shall be less than or 

equal to the profile given in Figure 3.3-4 (TBC). 

§3.4.1 For each kind of lines (command, acquisitions ...), all the available lines may be 

used but the total number of lines used by the Payload shall be lower than 75% 

(TBC) of these 268 lines. 

§3.4.4.1 Time tagged packets are regularly scanned to check if their due date is arrived. 

When this occurs, the packet is dispatched exactly as in the direct dispatching way. 

The dispatching time accuracy is estimated to ±250 ms (TBC) for 1553 

commands and ± 125 ms (TBC) for discrete command. 

§3.4.6 The platform offers nominally payload surveillance at 1/32 Hz and 1/8 Hz. It may 

also provide surveillance at 1 Hz. If any, the number of these last surveillance 

shall be lower than 20 (TBC). 

§3.5.3.1 During launch phase, the power lines 8 and 16 can be configured before launch, 

to supply power to part of the payload if needed (TBC depending on launch 

phase). But in this case they are not controlled by software before separation due 

to the OFF status of the data handling unit processors. 

§3.6.6 If these two ground braids can't be connected with the Payload Grounding Point 

(i.e. 2 ground studs as indicated on § 4.2.2.2), the payload supplier shall foresee 

2 dedicated ground studs as shown on Figure 3.6-8 (TBC values are typical values 

which shall be defined by the Payload Supplier depending on payload design). 

The ability to use Payload Grounding Point for STA grounding can be discussed 

with the Satellite Contractor. 




LIST OF TBDs 


Resolution 

§3.3.2 The power peaks demands of the payload during eclipse shall be lower TBD W (with 

TBD lower than 900 W) during TBD min. TBD are mission dependent. 

§3.3.2 The power peaks demands of the payload shall be lower than TBD W (with TBD lower 

than 900 W) during TBD min. TBD are mission dependent. 




3. PAYLOAD INTERFACE REQUIREMENTS 

This chapter deals with the mechanical, thermal, electrical, command & control interface requirements which must 

allow the payload to be compatible with the standard PROTEUS platform. An overview of the interfaces between 

payload and platform is given through these subjects. The information contained in the present chapter allows to 

design the payload at system level. The next chapters 4, 5 and 6 detail the payload other requirements; they are 

useful for the following step of the mission study. 

Note that, except opposite mention, all the payload requirements are expressed at equipped payload level that is to 

say «payload + STA + H02 & H03 brackets + STR cables (see section 1.5). 

STR cables shall be routed on the payload (see section 3.6.3) 

3.1 MECHANICAL INTERFACE REQUIREMENTS 

The platform/payload mechanical interfaces consist in: 

four surfaces which allow to mate the payload on the platform (see section 3.1.4.1) 

two electrical interface brackets (H02 & H03) which shall be accommodated on the payload (see section 

3.1.4.2) 

another electrical interface bracket (H01) which is located on the platform (see section 3.1.4.2) 

a Star Trackers Assembly (STA) which consists in 2 Star Trackers, their brackets and radiator and which shall 

be accommodated on the payload (see section 3.1.4.3). 

STR cables shall be routed on the payload (see section 3.6.3) 




3.1.1 MASS PROPERTIES 

3.1.1.1 Mass 

PL - 3.1.1 - 1 a 

The maximum allocated mass for the equipped payload is 300 kg. 

Mass properties given in the payload IDS shall be equipped payload mass properties (see appendix A). 

In order to estimate the satellite mass, the Table 3.1-1 gives PROTEUS masses. 

Mass (kg) 

Payload M – 14 

Payload balancing 0 * 

[STA + H02, H03 electrical brackets + STR wires] 14 ** 

Equipped payload Total M 

PROTEUS dry platform without [STA, H02, H03 brackets, STR wires] 269 

Platform balancing 0 * 

Dry Platform without [STA, H02, H03 brackets, STR wires] Total 269 

Launch vehicle adapter 15 *** (TBC) 

Hydrazine (maximum capacity of the tank) 28 

Satellite balancing mass 20 * 

Satellite maximum mass 332+ M 

* The balancing mass is a satellite one (no balancing mass is required at payload level because a system approach 

is preferred) which is limited to 20 kg and depends on the natural balancing of the payload. The rough 

determination of the needed balancing mass is obtained by Figure 3.1-2. 

** Considering the STA definition, the STA position on the payload and the STR cables routing on the payload are 

mission dependent, this maximum mass is an allocation. The maximum true mass shall be estimated when precised 

definitions will be known. 

*** Launch vehicle dependent 

Table 3.1-1: Satellite module masses 

3.1.1.2 Centering 

PL - 3.1.1 - 2 

The distance between the equipped payload (including STA and IF brackets) center of gravity location and 

the payload/platform interface plane must be lower than 0,73 m along the X P axis in launch and on orbit 

configuration. 

The payload/platform interface is defined in section 3.1.4. 



PRO.LB.0.NT.003.ASC 

PL - 3.1.1 - 3 a 

Issue. 06 rev. 03 Page: 3.19 

The equipped payload balancing (in the YZ plane), in launch and on orbit configurations, shall be 

compatible with a position of the satellite center of gravity inside a cylinder of radius equal to 5 mm around 

the satellite Xs axis. 

In order to express this requirement at equipped payload level, Figure 3.1-2 describe the allowed position of 

the equipped payload centre of gravity to get a balanced satellite versus counterbalance mass. Y and Z 

counterbalance masses shall be considered independently and the sum of these two masses shall be less 

than 20 kg. 

These counterbalance masses will be dispatched in the PF by ALCATEL SPACE. 

PL - 3.1.1 - 4 a 

The Payload Supplier shall determine the location of the CoG of the complete flight model of the payload to 

an accuracy better than 5 mm spherical error, in launch and deployed configurations. 

Figure 3.1-1 gives the equipped payload center of gravity (CoG) allowed Xp location in the Payload Reference Frame 

Fp with the corresponding satellite CoG Xs location in the Satellite Reference Frame Fs for different equipped 

payload masses. The platform mass is the maximum one, counted with hydrazine and launch vehicle adapter as 

described in Table 3.1-1, but without the balancing mass. 

Satellite CoG Xs location (mm) 

1200,0 

1100,0 

1000,0 

900,0 

800,0 

700,0 

600,0 

0 50 100 150 200 250 300 350 400 450 500 550 600 650 700 750 

Equipped payload CoG Xp location (mm) 



Equipped 

payload 

mass 

Figure 3.1-1 : Satellite CoG Xs location for different equipped payload masses and for different 

equipped payload CoG Xp locations 

Figure 3.1-2 gives : 

the envelope of the equipped payload CoG allowed Zp (Yp) locations in function of the equipped payload 

mass ; this envelope is defined by the maximal balancing mass, 20 kg, placed either on Zs = -460 (Ys = - 

460) or on Zs = 460 (Ys = 460) ; and 

100 kg 

150 kg 

200 kg 

250 kg 

300 kg


the envelope of the equipped payload CoG allowed Zp (Yp) locations in function of the equipped payload 

mass ; this envelope is defined by two maximal balancing masses limited to 10 kg, placed each one on a 

single point either on Zs = -460 (Ys = -460) or on Zs = 460 (Ys = 460) (Y and Z counterbalance masses 

shall be considered independently and the sum of these two masses shall be less than 20 kg); and 

the needed balancing mass and its Zs (Ys) location in order to maintain the satellite CoG Zs (Ys) location 

within the limits fixed by the requirement PL - 3.1.1 - 3. 

The platform mass is the maximum one, counted with hydrazine and launch vehicle adapter as described in Table 

3.1-1. 

-- 

PL centering allocation (Yp,Zp) 

40 

20 

0 

Yp 

(mm) 

-80 -60 -40 -20 0 20 40 60 

-20 

-40 

-60 

-80 

Zp 

(mm) 



PL 300 kg 

PL 250 kg 

PL 200 kg 

Figure 3.1-2 : Equipped payload CoG allowed Yp/Zp locations in function of the equipped payload 

mass


3.1.1.3 Inertia 


PL - 3.1.1 - 5 

The Payload Supplier shall determine each equipped payload principal inertia to an accuracy better than 5% 

of the total inertia, in launch and deployed configurations. 

3.1.1.3.1 In launch configuration 

PL - 3.1.1 - 6 

In launch configuration, the absolute values of the equipped payload inertia expressed in the Payload 

Reference Frame (O P, X P, Y P, Z P) F p shall be less than 75 kg.m 2 relative to the (O P, X P) axis and less than 300 

kg.m 2 (TBC) relative to the two other axes (O P, Y P) and (O P, Z P). 

PL - 3.1.1 - 7 

In launch configuration, the absolute values of the equipped payload crossed inertia expressed in the 

Payload Reference Frame (O P, X P, Y P, Z P) F p shall be less than 5 kg.m 2 . 

3.1.1.3.2 On orbit configuration 

PL - 3.1.1 - 8 

For on orbit configuration and for payload with deployable appendages, inertia limits will be discussed with 

the Satellite Contractor (orbit dependent). 

PL - 3.1.1 - 9 

For on orbit configuration, the difference between the equipped payload highest inertia expressed in the 

Payload Centre Of Gravity Reference Frame (Payload Reference Frame translated to the equipped payload 

CoG) and the equipped payload inertia relative to the (OP, XP) axis shall be lower than 50 kg.m². 

That is to say that : Imax – Ixx < 50 kg.m2 PL - 3.1.1 - 10 

For on orbit configuration, the absolute values of the equipped payload crossed inertia expressed in the 

Payload Reference Frame (O P, X P, Y P, Z P) F p shall be less than 5 kg.m 2 . 




3.1.2 STIFFNESS 

3.1.2.1 Stiffness in launch configuration 

PL - 3.1.2 - 1 

The minimum first mode frequencies allowed at equipped payload level in launch configuration are shown 

on Figure 3.1-4. 

The boundary conditions are hard mounted conditions on an infinitely rigid interface. 

Frequency [Hz] 


47,5 

45 

42,5 

50 

45 

40 

35 

30 

25 

50 

40 

Longitudinal Stiffness requirement for the Payload 

50 100 150 200 250 300 

Mass [kg] 

Lateral Stiffness requirement for the Payload 

50 100 150 200 250 300 

Mass [kg] 

Figure 3.1-4 : Stiffness requirements at equipped payload level 




3.1.2.2 Stiffness in flight configuration 


PL - 3.1.2 - 2 

The first mode frequencies allowed at equipped payload level in flight configuration shall be higher than 5 

Hz (TBC). 

Low frequencies domain shall be subject to specific mission analysis. 




3.1.3 AVAILABLE VOLUME FOR THE PAYLOAD 

3.1.3.1 Launch vehicles fairing constraint 

The volume allocated for the PROTEUS satellite by the main launch vehicle fairings and the satellite container is 

shown, for information, on Figure 3.1-5. Adaptation to smaller fairings is optional. 

PL - 3.1.3 - 1 

The envelope volume allocated to the equipped payload is the most constraining volume between the 

container volume and the fairing of the chosen launch vehicle. 

* the container volume envelope is characterised by a diameter of 2500 mm on a height of 3424 mm, this diameter 

is negotiable until 3280 mm (following the payload shape). 

Figure 3.1-5 : Payload volume under fairings 




PL - 3.1.3 - 2 


In order to accommodate the payload on the platform in the allocated volume, the following data shall be 

considered : 

• the height (along Xs axis) of the platform which is 1070 mm (show section 3.1.4.1) 

• the additional height of the launch vehicle adapter (example : 80 mm in case of DELTA II launch) type 

3.1.3.2 In flight configuration constraint 

In flight configuration, the volume dedicated to the payload is located above the platform pods and is limited first by 

the volume occupied by the solar array rotation around Y axis. Moreover, the payload shall not shadow this SA and 

the platform thermal radiators. 

PL - 3.1.3 - 3 

In flight configuration, the payload (excluding STA and local appendices) shall not exceed the volume 

described in Figure 3.1-6. 

Otherwise, the payload lay out on the platform shall be studied case by case (the critical points are mainly the 

shadow on the solar arrays, the payload and star trackers fields of view and the platform thermal control) and the 

Payload Supplier shall contact ALCATEL SPACE or CNES. 

For information, the volume occupied by the solar array is shown on Figure 3.1-7 and is described as follows : at 

815.5 mm from the +/-Y platform panels, the solar arrays can occupy a cylinder volume of 6.15 m3 of which the 

part interfering with the payload volume corresponds to a tunnel volume of height h = 550mm on a length L = 

3530 mm along Y axis for each side. 

45° 

270 mm 

170 mm 

PAYLOAD volume 

PROTEUS 

Platform 

955 mm 

IF platform/payload 



Xs 

Ys or Zs 

Figure 3.1-6 : Payload allowed volume in flight configuration


Figure 3.1-7 : Volume occupied by the solar array rotation 




3.1.4 MECHANICAL INTERFACES PAYLOAD/PLATFORM 

3.1.4.1 External interfaces for the payload accommodation 

3.1.4.1.1 Description 

Interface with the payload on the +Xs platform panel consists in four mechanical links at the four upper faces of four 

pods (cf. Figures 3.1-8 and 3.1-9). These pods are screwed on each upper corner fitting of the platform (cf. Figure 

3.1-10, Figure 3.1-11). 

This interface allows easy mounting and dismounting of the payload without opening neither the payload, nor the 

platform. It provides also a good thermal decoupling between platform and payload. 

The mating of the payload will be performed with the Xs mounting plane in a horizontal position. 

PL - 3.1.4 - 1 

The payload interfaces shall be compatible of those described Fig 3.1-8 to 3.1-11. 

Figure 3.1-8 : Platform Y view 

Figure 3.1-19 : Platform Z view 



.. 


Figure 3.1-9 : +Xs panel platform with its four PF/PL mechanical interfaces 




Figure 3.1-10: Payload fitting interface 

Figure 3.1-11 : Platform / payload interface details 




Figure 3.1-12 : Payload mounting interface detail, case of payload with the same structure as 

platform. 




3.1.4.1.2 Mounting surfaces 

The platform will provide four planar surfaces made of Titanium to which the payload is linked. 

Payload mounting bolts (with torquing tool) shall be provided by the Payload Supplier (see section 10.3 for delivery 

responsibility). 

PL - 3.1.4 - 2 

Platform/ payload interface sizing shall be under Payload Supplier responsibility. 

3.1.4.1.2.1 Flatness 

The relative flatness between the four mounting surfaces provided by the platform will be better than 0.1 mm 

(obtained by manufacturing or wedging) (TBC). 

PL - 3.1.4 - 3 a 

The surface defined by the four payload mounting faces shall have global flatness of 0.1 mm. 

3.1.4.1.2.2 Parallelism 

Each PF mounting surface will be parallel to the others with a precision better than 0.5 mrad. 

PL - 3.1.4 - 4 a 

The plane defined by the four payload mounting faces planes shall be parallel to the reference plane [Yp, 

Zp] with an accuracy of 0.35 mm. 

3.1.4.1.2.3 Roughness 

The PF mounting surface will have a surface roughness better than 3.2 micro m.rms. 

PL - 3.1.4 - 5 

The surface roughness of the payload mounting surfaces shall be better than 3.2 micro m.rms. 

3.1.4.1.2.4 Interfaces mechanical distortions 

The interface distortions will be defined specifically for each mission. 

PL - 3.1.4 - 6 

Nonetheless, the payload shall achieve its full performance with the following interface distortions coming 

from the platform. 

Origin of the distorions Type Values 

Orbital Flatness mission dependant value 

Parallelism mission dependant value 

Integration Flatness 0,2 mm (TBC) 

Parallelism 0,5 mrad 

Table 3.1-2 : Interface mechanical distortions 




3.1.4.2 Connector brackets interfaces 



The connector interfaces between payload and platform are distributed among three standard electrical interface 

brackets. The responsibility in the delivering items (brackets, connectors, harness, interface bolts) is given in section 

10.3. 

The « power » interface bracket (H01) is located in +Zs PROTEUS platform panel. The H01 connector bracket 

description is shown Figure 3.1-13. 

F 

Figure 3.1-13 : H01 connector bracket mechanical interface 

Nota: If payload does not use pyro, J03 and J06 places on H01 bracket stay empty 

F View 




PL - 3.1.4 - 7 


The two « acquisition and command » interface brackets (H02 and H03) shall be accommodated on the 

payload. 

Figure 3.1-14 gives the position and the size of the three interface brackets (JASON-1 example, mission dependant). 

Figure 3.1-15 shows these brackets global volume and Figure 3.1-16 gives their interface plane. 

Finally, the connector brackets mechanical description is given in Figure 3.1-17. 

The maximum mass of each bracket H02 and H03 is 0.15 Kg (this maximum mass is an allocation mass to take into 

account on the equipped payload mass calculation (see PL-3.1.1-1 specification). The maximum real mass shall be 

estimated when the lines will be affected and bracket definition known. 

These connector brackets electrical description (pin allocation) is given in section 3.5.2. 

Figure 3.1-14 : Electrical interface brackets global view 



-- 


Figure 3.1-15 : Electrical brackets volume 







Figure 3.1-16 : H02 & H03 electrical brackets : mechanical interface plane 

Figure 3.1-17 : H02 & H03 electrical brackets : connectors interface 

Warning: H02 & H03 electrical brackets mechanical interface plane and connectors interface will be updated if the 

J03 connector (carrying the CS16 lines) needs to be used. 




3.1.4.2.2 Connector Location 

PL - 3.1.4 - 8 

H02 and H03 Connectors shall be preferably located near the position shown in Figure 3.1-13 and 3.1-14. 

Moreover, a volume of ±120 mm around these brackets shall be reserved in order to be able to connect and 

disconnect connectors. 

Nonetheless, the Payload Supplier shall contact ALCATEL SPACE or CNES to discuss the real location of 

these connectors. 

PL - 3.1.4 - 9 

It shall be possible to grasp firmly the connectors for mating and demating. 




3.1.4.3 Star Trackers Assembly interfaces 


See section 3.6. 

3.1.4.3.2 Mounting surfaces 

3.1.4.3.2.1 Stability 

The position of the STA interface plane with respect to the platform / payload interface plane shall comply with the 2 

following requirements. 

PL - 3.1.4 - 10 

The thermoelastics effects between STA interface plane and payload interface plane shall be lower than 64 

arcsec, 

PL - 3.1.4 - 11 

Biases (including launch shift, gravity release and hygroelastic) between STA interface plane and payload 

interface plane shall be lower than 32 arcsec. 

These requirements could be negotiated according to mission pointing requirements. 

3.1.4.3.2.2 Flatness and parallelism 

PL - 3.1.4 - 12 

The parallelism between all the STA mounting surfaces shall be better than 0.4 mm for 400 mm. 

PL - 3.1.4 - 13 

The global flatness of each mounting surface shall be better than 0.1 mm 

Nota : See Figure 3.6.2b 




3.1.5 MAXIMUM GENERATED DISTURBANCES 

3.1.5.1 Dynamic disturbances 

In case of generation of such dynamic disturbances, the Payload Supplier shall contact ALCATEL SPACE or CNES and 

provide, for a first iteration, this kind of data : 

Total forces and moments at the PF/PL interface as a frequency spectrum 

Inertia of moving parts 

Kinetic momentum of turning parts 

Static and dynamic balancing 

Then, after this first iteration, the requirement could be the following : 

PL - 3.1.5 - 0 

Payload permanent dynamic disturbances are mission dependent and shall only be accepted after system 

analysis. 

PL - 3.1.5 - 1 

The permanent kinetic momentum due to a payload turning part at satellite level shall be lower than 

»mission dependent value». 

PL - 3.1.5 - 2 

The static balancing of the payload turning part shall be lower than »mission dependent value». 

PL - 3.1.5 - 3 

The dynamic balancing of the payload turning part shall be lower than »mission dependent value». 

3.1.5.2 Maximum Shock generated by the payload 

PL - 3.1.5 - 4 

The Payload shall not generate, at the platform / payload interface plane, a shock which leads to a shock 

response spectrum higher than the one given in Table 3.1-3. 

Frequency 

qualification level 

(Hz) 

(g) 

100 5 

2000 900 

10000 900 

Table 3.1-3 : Maximum Shock levels generated at the PF/PL I/F plane by the Payload 

3.1.6 OPTIONAL PAYLOAD MODULE 

As an extended PROTEUS service, a standard payload module design, adaptable in height, mechanically qualified, is 

proposed : 

either to accommodate a payload made of several boxes, 




either to be an intermediate module between the platform and a main instrument (for example a telescope..) 

to contain various electronics and/or an optional X band data communication subsystem (mass memory, 

modulators, amplifiers, switch). 

Figure 3.1-18 : Payload JASON module as example 

The payload module as shown on Figure 3.1-18 presents the same structure as the platform one: it is a cubic shape 

with no central structure, the panels making the cube have both functions of providing structural strength as well as 

surface to accommodate equipment. Lateral panels provide heat rejection surfaces for thermal control of the 

module. Interface with the platform is provided through the four upper corners of the platform, with the pods out of 

titanium alloy. The interface with the payload is realized in the same way. In fact, the payload module presents four 

mechanical links at the four upper corners as the platform ones (see section 3.1.4). Satellite design is thought such 

that each module (platform, payload module, payload) is thermally decoupled. 




3.2 THERMAL INTERFACE REQUIREMENTS 

The purpose of the satellite thermal control is to maintain all the elements of a satellite system within their 

temperature limits for all mission phases including safe hold mode. For the launching phase, active thermal control 

is not foreseen. 




3.2.1 PLATFORM-PAYLOAD CONDUCTIVE AND RADIATIVE INTERFACES 

PL - 3.2.1 - 1 

The payload is thermally decoupled from the PROTEUS platform through the four interface pods made of 

titanium alloy. A MLI blanket is accommodated between the PROTEUS platform and the payload module on 

the +Xs panel. The equivalent efficiency of this blanket is 0.1 W/m²/°C. Moreover, an ALCATEL provided MLI 

skirt is also accommodated as shown in Figure 3.2-3. The remaining thermal conductive coupling can be 

assumed to be less than 0.04 W/°C for each titanium pod. 

The Payload Supplier shall size its thermal control assuming this remaining conductive coupling and 

assuming a platform structure temperature between -5 °C and +40 °C. 

PL - 3.2.1 - 2 

The Payload Supplier shall size its thermal control assuming the radiative coupling based on the following 

assumptions relating to the platform surfaces : 

• SSM areas between -25°C and +40°C 

• MLI areas in adiabatic equilibrium with the environment 

• Solar Array between -100°C and +95°C 

The thermo-optic characteristics of the platform are given in Table 3.2-1. 

ε 

InfraRed Emissivity 

α min 

Solar Absorptivity 



α max 

Solar Absorptivity 

SSM 0.76 0.10 0.16 

MLI 0.77 0.32 0.49 

Solar array (Solar cells face) 0.82 0.75 0.85 

Solar array (back face) 0.7 0.92 0.92 

Table 3.2-1 : Thermo-optic characteristics of the platform parts 

MLI typical interface between platform and payload is shown on Figure 3.2-3. 

Typical sizes of SSM surfaces are given in Figure 3.2-1. The other part of the panels may be considered as MLI 

surfaces. 

The Solar Array dimensions are shown on Figure 3.1-7. 

Dimensions of the Platform as indicated in Figures 3.1-7 and 3.2-1 may differ due to MLI thickness.


PAYLOAD 

Payload 

5 mm 

Platform 

5 mm 

5 mm 

L 

INTERFACE MLI 

A 

A 

80 mm < L < 150 mm 

TYPICAL VELCRO (25 x 25) 

A - A 

INTERFACE MLI 

PLATFORM 

PAYLOAD 

PLATFORM 

Figure 3.2-3 : MLI typical interface between platform and payload 



VELCRO ASTRAKAN 

VELCRO HOOK 

INTERFACE MLI


874 

874 

140 

100 

165 

124 

625 

955 

706 

955 

594 

674 

Figure 3.2-1 : Typical dimensions of Platform thermal radiators 



125 

124 

706 

706 

955 

955 

674 

674 

100 

100


3.2.2 ACTIVE THERMAL CONTROL 

As the payload is completely decoupled from the platform, it can be in charge of its own thermal control, no 

constraint by the platform is imposed at payload level. 

If the payload requires an active thermal control, then two options are offered: 

Option 1: the payload is in charge of its own thermal control and it uses the power lines described in section 3.5.3. 

This is possible because the payload is decoupled from the platform. But it should be noticed that in this case and as 

described in section 3.5.3, no power will be available during SHM phases and that therefore no active 

thermal control will be available during SHM phases. 

Option 2: the payload uses the heating lines provided by the platform and described in section 3.2.2.1. These 

heating lines are continuously active, except during the switching phase to SHM when reconfiguration 

occurs (typically 1 minute). They are controlled by the OBSW starting from the launcher separation (see section 

3.2.2.2). In case of reconfiguration by FDIR, an automatic swap from nominal to redundant (or vice versa) will be 

performed. 

These two options can be mixed. 

3.2.2.1 Heaters lines and thermistors 

PL - 3.2.2 - 1 

A maximum of eleven heating lines (11 nominal and 11 redundant) shall be used by the payload. Each line 

is associated to three temperature acquisition sensors which must be located inside a circle of 15 mm of 

diameter. 

PL - 3.2.2 -2 

The maximum power available for these 11 lines is distributed in the following way : 

• 3 lines of 50 W under 28 V (lines number 17, 18 and 21), 

• 4 lines of 25 W under 28 V (lines number 15, 16, 19 and 20), 

• 4 lines of 10 W under 28 V (lines number 01 to 04). 

PL - 3.2.2 - 3 

Voltage applied on heaters will be provided by the BNR (23-37 V). 

Note : a heating line is able to provide any heating power lower than its design value (i.e. a 25 W line may provide 

8 W only if needed in case all the 10 W lines are allocated). 

Nominal heaters are controlled by DHU/PM A (nominal processor module). 

Redundant heaters are controlled by DHU/PM B (redundant processor module). 

The 3 thermistors (for each heaters line) monitor both nominal and redundant heaters. 

Thermistors are shared components, controlled by PMA and PMB (cf. Figure 3.2-2). 




Figure 3.2-2 : Heaters and thermistors redundancy configuration for thermal control 

PL - 3.2.2 - 4 

Heater lines are considered as pure resistors. They shall be dimensioned at an average voltage of 28 V. 

PL - 3.2.2 - 5 

Thermistors type shall be Fenwal Fw 526-31 BS12-153. 




3.2.2.2 Regulation algorithm 


The heating lines are controlled in closed loop (PI algorithm) by a regulation algorithm running on the PROTEUS On 

Board Software. 

In this algorithm, the heating lines are described by : 

one line identifier, 

one regulation authorization flag, 

one heater command address, 

three thermistor acquisition addresses, 

the maximum power dissipated by the heater P Heat under the V Heat command: 

P Heatt is to be given by the Payload Supplier for each line. 

V Heat is a constant value set to V Heat = 28 V in the Satellite database. 

the target temperature T t, 

the PI correction coefficients C1 and C2. The standard thermal control works at 1/32 Hz with a commands resolution of 1 s. 

The thermal control loop executes at each cycle (identified by k), the following steps : 

Get the 3 thermistors measurements{T k_Th1, T k_Th2, T k_Th3}, 

Compute T k as the median value within the triplet {T k_Th1, T k_Th2, T k_Th3}. 

Compute the power injection command P k(T t, T k, P k-1), 

Apply Pk command algorithm. 

The power injection command Pk is computed as follows : 

Compute the temperature error : E k = -1 * (T t - T k) 

Compute the PI correction : S k = P k-1 + C 1*E k + C 2*E k-1 

Compute power injection command : 

if (0 < S k 

else if (S k > P heat) then P k = P heat 

else P k = 0 

Memorize Ek and Pk for next regulation cycle 

The power injection command computation algorithm is initialized for each control loop as follows : 

P 0 = 0 

E 0 = 0 

k = 1 

On ground and for test purpose, the thermal control SW application could be deactivated or each line could be 

commanded independently with a constant power injection command (for thermal balance test or other test). 

The Pk command algorithm consists in : 

Computing the Heating Duration within the 32 seconds regulation cycle, assuming that the Heating Duration 

granularity is 1 second and that the power dissipated by heater is an instantaneous power (P max) dependant on 

the voltage : 




Pmax = (VDHU / Vheat) Issue. 06 rev. 03 Page: 3.48 

square * Pheat Heating Duration = Min {E (Pk * 32 / Pmax), 32}, where E is the closest integer value. 

TH application will continuously overload the computed Heating Duration, if it is authorized, with the uploaded value 

for the associated line. 

For each line, the default initial value of the Heating Duration overloading is inhibited. 

TH application shall apply the Heating Duration (computed or overloaded) as follow: 

Start heating (heater ON command), if Heating Duration is different from 0, immediately at the beginning of 

the next one second regulation cycle, 

Stop heating (heater OFF command), if Heating Duration is different from 0, when Heating Duration is 

expired. 

At each begin of thermal control loop, the commands heater OFF of previous thermal control loops shall be sent 

prior to the eventual command heater ON. 

The default initial values of C1, C2, Tt and Vheat will be extracted from satellite database at OBSW generation. 

Thermal control set of adjustment parameters consist in all thermal lines parameters : 

Target temperature T t, 

Coefficients C1 and C2. These parameters are modifiable by telecommand. 

PL - 3.2.2 - 6 

These thermal parameters (C 1, C 2, T t) shall be defined by the payload thermal control responsible for each 

line and each payload mode. 

PL - 3.2.2 -8 a 

These thermal parameters (C1, C2) shall be defined as follows: 

• C1 < 0 

• C 2 > 0 

• ⏐C 1⏐ > ⏐C 2⏐ 

• A difference between ⏐C 1⏐ and ⏐C 2⏐ higher than 0.05 W/°C is recommanded. 

PL - 3.2.2 - 7 

Electrical thermal sensors characteristics shall be compliant with requirements of section 3.5.6.2.2. 




3.2.3 PAYLOAD THERMAL MONITORING 

In addition to the 11 heating lines and the 33 associated acquisitions lines offered for active thermal control, the 

Satellite Contractor provides to the Payload up to 48 temperature acquisition lines for thermal monitoring (see 

section 3.4.1). These lines are split in 3 independant groups as described in section 3.4.1. 

The thermal sensors compatible with these acquisition lines are the Fenwal Fw 526-31 BS12-153 var 32 and the 

Rosemount 118 MF. 

PL - 3.2.3 - 1 

The respective number of each sensor shall be lower than : 

• 36 Fenwal Fw 526-31 BS12-153 

• 12 Rosemount 118 MF 2000 

If necessary, this standard repartition shall be negotiated with ALCATEL SPACE or CNES. 

These thermal sensors will be powered by the satellite and conditioned in the satellite data handling subsystem. They 

will be read out in the satellite housekeeping data stream when the payload is either ON or OFF except for the 

launch phase and for the SHM until the first acquisition of the satellite (cf. section 3.4.6). 

The Fenwal thermistor measurement range is from -60 °C to +90 °C. 

The Rosemount sensor measurement range is from -120 °C to +140°C. 




3.3 POWER SUPPLY INTERFACE REQUIREMENTS 

3.3.1 MEAN ORBITAL POWER AVAILABLE FOR THE PAYLOAD 

On each orbit, and for each pointing type, the maximum available mean orbital power for the payload can be 

calculated, searching the limit when the battery Deep Of Discharge returns to Zero just before eclipse entry in the 

worse case (End Of Life, worst case seasonal effect and maximum Solar Array pointing error in the case of yaw 

steering), assuming a constant power consumption. 

The calculation (cf. Figure 3.3-1) gives the total satellite available power on the PROTEUS flight domain for four solar 

array strings failed after a life duration of 3 years. The calculation results with the previous hypothesis (3 years and 4 

lost strings) are similar to the ones with 5 years and no platform failure. So in a first approach, these estimations can 

be used to lead studies about PROTEUS based missions with a life duration of 5 years. The same calculation (cf. 

Figure 3.3-2) is done in the particular case of sun synchronous orbits, assuming a perfect 3 axis pointing toward the 

Earth (no yaw). This estimation is given at the worst case date. The curve is not symmetrical around noon due to Sun 

declination. For orbits far from12 hours, the power loss on the solar array (cosine effect due to solar array angle 

versus Sun direction) is partially compensated by a better battery charge/discharge efficiency : different SA 

temperature, different SA voltage, different battery charge and a reduced eclipse duration. 

Then, the same simulation (cf. Figure 3.3-3) is done but for sun synchronous orbits 6h00-18h00, assuming a perfect 

3 axis pointing toward the Earth (no yaw) and considering two cases : the solar arrays are fixed and the solar arrays 

turn around their axis (Ys) to optimise the Sun incidence on the solar arrays. This estimation is given on the 21th June 

with an ascending node of 18h00 corresponding to a solar flux close to the minimum value and a maximum eclipse 

duration. The User can distinguish three areas for these curves : 

the first one with an altitude varying from 500 to around 1350 km corresponds to an orbit of which one part is 

in total eclipse ; the satellite power increases as the eclipse duration decreases. 

the second one from 1350 km to 1400 km corresponds to an orbit of which one part is in penumbra ; that 

explains the satellite power increasing. 

the third one from 1400 km to 1500 km corresponds to an orbit with no eclipse. In this area, the radiation 

effects perturb the solar array power gain. 

PL - 3.3.1 - 1 

The maximum mean power available for the payload per orbit is given by these abaci* assuming the power 

drained by the platform is typically 300 W. 

Moreover, the mean consumption of the payload during eclipse shall be lower than or equal to the mean 

orbital power. 

Notice : The power consumed for the payload thermal control must be taken into account in the payload 

power budget. 

* The following abaci are estimated with Jason electrical architecture hypothesis; for the Standard Proteus platform, 

the values shown in these abaci are guaranteed as minimum in all cases and will be updated in the next PUM 

edition. 

Any other payload consumption profile (average on several orbits for example) may be discussed with ALCATEL 

SPACE and CNES. 




Figure 3.3-1 : Maximal satellite mean orbital power w.r.t the altitude and the inclination 

(EOL 3 years, 21/06, sun at 15 deg of the orbital plane without Yaw steering , 4 lost string, 

ascending node 12 h, Battery temp = 10°C) 

Figure 3.3-2 : Maximal satellite mean orbital power w.r.t. the ascending node of the sun 

synchronous orbit and the altitude 

(EOL 3 years, worst case date, 4 lost string, battery temp. = 10°C) 




(Y coordinate represents mean local hour; that is to say instantaneous local hour varies around this mean value of 

±16 min according to the time equation) 

Figure 3.3-3 : Maximal satellite mean orbital power versus the altitude for Sun-synchronous orbits 

LHAN 18 h 

Payload power = Satellite power - Platform power (300 W) 

(EOL 3 years, summer solstice, 4 lost string) 

3.3.2 POWER PEAKS LIMITATIONS FOR THE PAYLOAD 

In addition with the respect of the mean orbital power available (specification PL - 3.3.1 - 1), the Payload shall 

comply with the following requirements. 

PL - 3.3.2 - 1 

The power peaks demands of the payload during eclipse shall be lower TBD W (with TBD lower than 900 W) 

during TBD min. TBD are mission dependent. 

PL - 3.3.2 - 2 

The power peaks demands of the payload shall be lower than TBD W (with TBD lower than 900 W) during 

TBD min. TBD are mission dependent. 




3.3.3 POWER SUPPLY DURING TRANSIENTS 

3.3.3.1 Launch phase 

Nominally, no power is delivered to the payload. 

3.3.3.2 SHM phase 

PL – 3.3.3 - 1 

Payload power consumption (including thermal control) shall be less than or equal to the profile given in 

Figure 3.3-4 (TBC). 

Power (W) 

140 

120 

100 

80 

60 

40 

20 

0 

T 0 T 1 

2 4 6 8 10 12 

Duration (h) 

T1 = 8 minutes for first SHM (after separation from the launch vehicle) 

T1 = 1 minute for other SHM 

Figure 3.3-4 : Allocation for payload power consumption during SHM phase 

3.3.3.3 Orbit Control phase 

PL – 3.3.3 - 2 

Payload power consumption (including thermal control) during orbit control phase shall be less than or 

equal to «mission dependent value». 




3.4 COMMAND & CONTROL INTERFACE REQUIREMENTS 

3.4.1 COMMAND AND CONTROL AVAILABLE LINES 

The standard PROTEUS platform Command and Control functional chain provides the following electrical interfaces 

to the payload: 

1 MIL-STD-1553B redundant bus (possibility to connect up to 12 units or instruments) 

command lines: 

16 pyrotechnic lines, protected by three safety barriers (redundancy has to be chosen among these 16 lines 

by the Payload Supplier), 

56 relay command lines, also named HLC for High Level Commands (power management excluded), 

10 CS16 (16 bits serial command lines), 

20 LLC (low level command lines), 

acquisition lines: 

28 relay status lines, also named DRS for Digital Relay Status, 

10 logic status acquisition lines, also named DB for Digital Bilevel 

16 AS16 (16 bits serial acquisition lines), 

56 analogous acquisition lines (ANA), 

48 temperature acquisition lines (acquisition lines dedicated to thermal control excluded), 

8 lines delivering simultaneously one pulse per second for datation. Information on date are given through the 

MIL-STD-1553B bus. 

PL - 3.4.1 - 1 

For each kind of lines (command, acquisitions ...), all the available lines may be used but the total number of 

lines used by the Payload shall be lower than 75% (TBC) of these 268 lines. 

The standard PROTEUS platform Command and Control resources are distributed as follows in three independent 

groups (P1, P2, P3) inside the DHU: 

HLC (*) LLC CS16 (**) DRS (*) DB (*) AS16 (**) ANA (*) TEMP 

P1 20 7 1(×2) +1(×1) 10 3 1(×3) + 1(×2) 18 16 

P2 16 6 2(×2) 8 4 2(×3) 20 16 

P3 20 7 1(×2) +1(×1) 10 3 1(×3) + 1(×2) 18 16 

Total 56 20 10, of which 6 

are independent 

28 10 16, of which 6 

are independent 



56 48 

Table 3.4-1: Distribution of the Command & Control resources inside the DHU 

Remark: The 3 independent groups (P1, P2 & P3) indicated on the Table above correspond with respectively the 3 

independent cards (SiOP 3, SiOP 4 and SiOP 5) implemented inside the DHU. 

(*) Note: 1 common return for 2 lines. 

PL - 3.4.1 - 2 

Two lines having the same return shall be allocated to the same unit.


(**) Note: CS16 and AS16 have multiplex capability. 

Notation n(×3) means n independent lines, each one having 3 enable signals. 

For each multiplexed line, the wiring is as follows: 

1 clock signal (2 wires), 

1 data signal (2 wires), 

up to 3 enable signals (2 wires per enable signal). 

PL - 3.4.1 - 3 

Using this capability, one multiplexed AS16 (with 3 enable signals) is equivalent to three AS16. All the enable 

signals of a multiplexed CS16 or AS16 shall be connected to the same unit. 

Some CC resources from P1 and P2 share the same connector on H02. Idem for P2 and P3 on H03. This could 

induce a deviation from standard SPF (Single Point Failure) rules . This point shall be discussed (mission dependant) 

with ALCATEL and CNES. 




3.4.2 PAYLOAD INSTRUMENT COMMAND/CONTROL STATUS 

Payload instruments are monitored by the platform using four status: 

ISOLATED, PASSIVE, STAND-BY STATUSES: The 1553 bus is not used. Even if these modes are equivalent 

from a functional point of view, they have been kept because on ground they represent different states of 

confidence in OBDH measures acquired on payload, and they can be used as a decommutation criterion. 

OPERATIONAL STAUS : The 1553 bus is used for the TC/TM transit between platform and payload. 

OPERATIONAL STATUS: The 1553 bus is used for the TC/TM transit between platform and payload units. 

Any transition between the instrument states is allowed from ground command. Instrument powering is performed by 

a specific telecommand (OBDH one) before the transition between passive state to stand by or operational state. 

All the transitions are performed under ground control and after operational coordination. 

The automatic transition performed by the OBSW are described into section 3.4.6.1. 

TC or PL 

anomaly 

ISOLATED 

TC 

PASSIVE 

TC TC 

TC 

STANDBY OPERATIONAL 



TC 

TC or 1553 anomaly 

TC or PL 

anomaly 

Figure 3.4-1 : Payload instrument command/control status


3.4.3 MIL-STD-1553B DATA BUS INTERFACE 

The DHU will implement all the standard features of the 1553B standard using the implementation of the DDC BU 

61582 bus controller. 

The DHU will operate as Bus Controller (BC), and the payload instruments as Remote Terminals (RTs). 

A protocol level (level 3) was built above the standard protocol (level 2) in order to allow the foreseen BC to RT and 

RT to BC exchanges. This protocol level 3 is described hereafter. 

A general view of 1553 BC to RT and RT to BC exchanges and associated timing is given hereafter. It shall be 

noticed that the proposed protocol is based on asynchronous communication mechanisms and that its timing will 

highly differ from one 250 ms cycle to another. Indeed, this timing is highly dependent on OBSW activities such as 

AOCS activities, telemetry acquisition and command dispatching which are not constant along the time. 

Consequently, figure 3.4-4 provides the description of the 1553 activities on one 250 ms cycle for two examples: 

the first one gives the timing of BC to RT and RT to BC exchanges in a typical case where a few PL data is 

transferred on the bus, 

the second one gives the same timing exchanges in a case where a lot of PL data is transferred on the bus. 

In addition, it shall be noticed that these 250 ms cycles are not synchronised with PPS delivery. 

Steps 1, 2, 3, 4, 5, 6 of this cycle are executed in chronological order. 

Figure 3.4-4: Timing of typical 1553 exchanges between platform and payload (corresponding 

respectively to few and many data transferred on the bus) 




3.4.3.1 System requirements 


PL - 3.4.3 - 1 a 

The payload units shall dialog with the central OBSW via a MIL-STD-1553B bus. 

A line may not be associated to Remote Terminal. 

PL - 3.4.3 - 20 

The MIL-STD-1553B interface between Payload units and H02/H03 connectors brackets shall comply with 

the Figure 3.4-6 configuration. 

PL - 3.4.3 - 2 

This bus will be compliant with the MIL-STD-1553B notice 2. 

The BC interface coupler is the DDC BU 61 582. 

PL - 3.4.3 - 3 

The length of the transmitted message from the RT to the BC shall be less or equal to 512 words. 

PL - 3.4.3 - 4 

The RT to BC messages shall be encoded using the CCSDS format. 



H02 


B 

AD 

Z2N 

PL 

PL Unit connected 

to 1553 

AD AD 

AD AD 

AD AD 

H02-P/J05 H03-P/J05 

Test point HO4 

J05 & J06 

DHU 

DHU-P044 DHU-P094 

P/F 

Z2R 

AD 

AD 

B Z1N 

Z1R B 

Bus N Bus R 



B 

PL 

H03 

Figure 3.4-6: interface between Payload Unit connected to 1553 and H02/H03 

Nota : the P/F test point H04 J05 & J06, allowing 1553 spying, is available along AIT satellite campaign for 

investigation in case of anomaly or any specific request. 

PL - 3.4.3 - 5 

Deleted. 

BC to RT messages will not be CCSDS encoded. 

PL - 3.4.3 - 6 a 

The system will offer the possibility to connect up to 16 units (each unit shall be shared, that is to say 

accessible by the nominal and redundant PMs of the DHU). 

P/F


Note that the total number of RT addresses is 30 and that the assignment of one address to one RT will be done 

through the IDS. 

3.4.3.2 Description of payload units behaviour 

PL - 3.4.3 - 7 

deleted 

As 1553 communication protocol is used between platform and payload units, payload units send messages in an 

asynchronous way. 

PL - 3.4.3 - 8 

deleted 

PL - 3.4.3 - 9 

Payload units shall change their mode autonomously (according to experiment conditions), or on ground 

request. Proteus platform does not manage internal modes of the payload units. 

PL - 3.4.3 - 10 

deleted 

PL - 3.4.3 - 11 

deleted 

3.4.3.3 Protocol general requirements 

PL - 3.4.3 - 12 

The length of a message shall be constant for a given message type. 

PL - 3.4.3 - 13 a 

Message types will be specific to a payload unit ; a message type shall correspond to a RT subaddress. 

PL - 3.4.3 - 14 

The knowledge of the different lengths associated with each message shall be communicated to the BC 

Software developers, using IDS, during the development phase. Then, they will be part of the satellite 

database. 

PL - 3.4.3 - 15 

Deleted 




PL - 3.4.3 - 16 


Priority between RTs for acquisitions is managed by the OBSW according to the following rule : 

• RT selection : priority decreases from RT address 1 to RT address 30 (ie RT@1 has the highest priority, 

RT@30 has the smallest priority). 

• Message selection : priority increases from message subaddress 1 to subaddress 14 (ie subaddress 1 

has the smallest priority, subaddress 14 has the highest priority). Subaddress 15 (reserved) shall 

correspond to MSB. 

For commands, messages date determines the order. 

3.4.3.4 Data Bus interface characteristics 

PL - 3.4.3 - 17 

The communications allowed are BC to RT, RT to BC, BC to all RTs. There is no RT to RT exchange. 

PL - 3.4.3 -21 

The instrument units shall comply with the following restrictions of 1553 standard features: 

1. Types of messages 

RT to RT transfer information is not used on PROTEUS. 

2. Status Word 

Here is a picture of the 1553 status word showing bit times (including synchronization and parity bits which are 

not useful bit for the payload). Useful bits are also numbered. 

Bit times 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 

Useful bits 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 

Sync 

Remote terminal 

Address 

Figure 3.4-5: 1553 status word 

Remote_terminal_address (5 bits:0 to 4) and message_error bit (bit5) are mandatory. 

Other optional status bit of the standard are used as follows: 

Message error bit 

Instrumentation 

bit 6: no use of the instrumentation bit, set to zero. 



Service request 

Reserved 

Broad cast command 

Busy bit 

Subsystem flag 

Dynamic bus control 

acceptance 

Terminal flag 

bit 7:the service_request bit will be used; refer to section 3.4.5.3.1 (RT to BC transfer protocol) 

bits 8, 9, 10: reserved 

bit 11: the broadcast_command bit will be used; it will be acquired with a transmit_status_word or 

transmit_last_command mode code message. 

bit 12: no use of the busy_bit; set to zero. 

bit 13: no use of the sub-system_flag bit; set to zero. 

bit 14: no use of the dynamic_bus_control bit; set to zero. 

bit 15: the terminal_flag_bit is not used by the OBSW. 

Parity bit


3. Command types 


The following mode_code commands will be used (only the white column has to be considered by the payload; the 

3 other columns are given for information): 

Mode 

Code 

Function Broadcast 

allowed 

(1553 standard) 

may be used at may be sent under 

Instrument level Ground Control or 

during Satellite AIT 



used by 

Platform 

OBSW 

00000 Dynamic bus control No No No No 

00001 Synchronise Yes No No No 

00010 Transmit status word No Yes No No 

00011 Initiate self test Yes Yes Yes No 

00100 Transmitter shutdown Yes No No No 

00101 Override transmitter shutdown Yes No No No 

00110 Inhibit terminal flag bit Yes Yes Yes No 

00111 Override inhibit terminal flag bit Yes Yes Yes No 

01000 Reset remote terminal Yes Yes Yes No 

01001 to 

01111 

Reserved 

10000 Transmit vector word No Yes No Yes 

10001 Synchronise with data word Yes Yes No Yes 

10010 Transmit last command No Yes Yes No 

10011 Transmit BIT word No Yes Yes No 

10100 Selected transmitter shutdown Yes No No No 

10101 Override selected transmitter shutdown Yes No No No 

10110 to 

11111 

Reserved 

Table 3.4-2: "mode_code" commands usable by the Payload 

Address 31(1FH) is reserved for broadcast command management. 

Subaddress 16 (10H) is reserved for broadcast reception within each RT. 

Subaddress 30 (1EH) is reserved for data wrap-around. 

Subaddresses 31 (1FH) and 0 (0H) are reserved to indicate a mode_code command. 

Address use Sub address use 

31d broadcast address 16d broadcast subaddress 

0d indication of a mode_code command 

30d reserved for data wrap-around capability 

indication of a mode_code command 

31 d 

Table 3.4-3: Subaddresses of the address 31 reserved for broadcast command management 

Some mode_code commands may be sent by Ground for investigation purpose (refer to Command Types table 

herebefore). 

The ones which do not require any RT answer will be sent to the RT using a standard 1553 TC service within 

the OBSW. 

The mode_code commands requiring a two word answer will be sent to the OBSW via a specific 

telecommand. 

When executing this telecommand, the OBSW will wait for the answer (status word + one word), then the 

OBSW will built an asynchronous TM packet to reply the RT answer to the Ground.


This packet will be found in the platform housekeeping telemetry (TMD, LTTM, FDTM). 

The concerned mode_code commands are the followings : 

TRANSMIT_LAST_COMMAND 

TRANSMIT_BUILT_IN_TEST_WORD 

Only one of these two mode_code commands will be transferred by the OBSW during a 250 ms cycle for the whole 

payload. 

3.4.3.5 Initialisation of the protocol 

PL - 3.4.3 - 18 

There is no initialisation phase of the 1553 level 3 protocol. 

The payload instruments will be turned on and managed by Ground until they are operational (observed via 

the AS16, ANA, DR status). 

The Ground Segment is able to manage 4 status of the instruments within the OBSW : 

isolated, 

passive, 

standby, 

operational. 

These status are managed via specific OBSW telecommand. 

The BC will not send a synchronize command at the initialisation. 

PL - 3.4.3 - 19 

RTs shall be able to receive commands to (if necessary) : 

• Load the SW of the RT Host 

• Change RT application mode (ex : calibration, standby, operational....). 




3.4.4 PAYLOAD COMMANDABILITY 


The orders issued from ground are received, decoded, stored, and sent by the platform DHU to the Payload 

Instrument. These orders are of several types : 

telecommands sent to the payload for immediate execution (TCUI) 

telecommands sent to the payload for delayed execution, called time tagged telecommands (TCUT) 

The time tagged telecommand service is characterized by the following rules : 

Time tagged packets are regularly scanned to check if their due date is arrived. When this occurs, the packet is 

dispatched exactly as in the direct dispatching way. The dispatching time accuracy is estimated to ±250 ms 

(TBC) for 1553 commands and ± 125 ms (TBC) for discrete command. 

Time tagged ground commands are dated using UTC. 

A TCUT command has priority on an immediate one. 

If the date of the TCUT command is older than 16s from the current time, it is rejected and a message is sent 

to ground. 

There is no TCUT command during the Safe Hold Mode. 

Ground is able to suppress TCUT commands in the TCT command file between two dates. This telecommand 

is an immediate one. 

Ground commands may be protected by using standard CCSDS mechanism called «authentication». This protection 

avoids any external intrusion during the satellite’s operation. 

Payload commands are distributed to the payload using discrete lines or the MIL-STD-1553 bus. 

PL - 3.4.4 - 1 

Deleted 

3.4.4.2 Discrete commands 

PL - 3.4.4 - 2 

Commands will be sent asynchronously towards the payload by the Platform On Board Software. 

PL - 3.4.4 - 3 

The Commands will be delivered to the instruments on time: due date will be managed by the software 

included in the Platform DHU (Data Handling Unit). 




3.4.4.3 1553 commands 


3.4.4.3.1 Telecommand protocol (BC to RT transfer) 

3.4.4.3.1.1 BC to RT Functional requirements 

PL - 3.4.4 - 4 

No particular protocol is foreseen to manage telecommand transfer from BC to RT. 

• The number of words within a telecommand is less or equal to 32 (1w=16 bits). 

• Each telecommand shall be contained in one 1553 message 

• The RT shall accept all types of commands that Ground can send to a particular unit. Those commands 

have to be included within the command/control IDS. 

Note : be careful when building 1553 commands : in the 1553 standard, a length equal to zero means 32 words. 

3.4.4.3.1.2 BC to RT timing requirements 

As well for immediate or time tagged commands, messages are sent keeping them in chronological order. 

PL - 3.4.4 - 5 

The BC will process and send the telecommands issued from ground not before 51 ms after the end of the 

polling sequence. 

An allocation of 32 ms is foreseen for that purpose on the bus. 

In addition, it shall be noticed that the platform will only provide a [8µs, 11µs] inter-message gap (1553 standard 

value). There are no other minimum or maximum timing constraints than the ones imposed by the standard protocol 

between the end of a message and the beginning of the next one. 

The polling sequence includes: 

The sending of the Transmit_Vector_Word towards the RT (step 1) 

The analysis of the Vector_Word transmitted by the RT (step 2). 

Figure 3.4-4 provided in section 3.4.3 illustrates this timing requirement. 

PL - 3.4.4 - 6 a 

As soon as at least one RT is in operational state, 

• Every second the RT shall accept a BROADCAST COMMAND (remote terminal address : 31 ie 1FH, Sub 

address 16d (see table 3.4-3) 

• this BROADCAST COMMAND includes the datation of the next pps (cf. Time and synchronisation 

section). 

Nota : As soon as a Payload unit is declared operational by OBSW (upon ground Tc), 1553 dialog starts 

between DHU and this Payload Unit. 

3.4.4.3.1.3 Error Handling on BC to RT transfer 

PL - 3.4.4 - 7 

1553 commands issued from Ground will be rejected if the concerned instrument is not known as « 

operational » by the OBSW. 




PL - 3.4.4 - 8 


In case of an anomaly detection (either non response or anomaly detected at BC level), the instrument on 

which the anomaly is detected will be put in STANDBY status by the OBSW (no longer polled regarding the 

1553 communications, but observable via the discrete acquisitions). 

Here are the possible 1553 errors : 

Time out on 1553 fault register 

Handshake error on 1553 fault register 

Format error on 1553 fault register 

Status word error on 1553 fault register 

Detailed explanation of these errors are given in the RT controller Data Sheet. 

3.4.4.3.1.4 BC to RT Protocol limitations 

PL - 3.4.4 - 12 

The number of types of messages (broadcast excepted) is limited to 14 max per RT (subaddresses 1 to 14). 

PL - 3.4.4 - 9 

The maximum number of TC messages of 32 words each RT shall be able to manage during a 250 ms cycle 

is 8 (peak level). 

The maximum peak command capacity is then 20,5 kbits/s (8×32×20 (16 useful bits + 1parity bit + 3 

synchronisation bits) ×4). 

However, average data rate to be considered for mission sizing purpose is mission dependent and is an output of 

system analysis taking into account ground-to-board constraints. 

GR - 3.4.4 - 1 

It is the responsibility of the Ground Segment to ensure that the rate is not overrun. 

The Commands will be delivered to the instruments on time: due date will be managed by the software included in 

the Platform DHU (Data Handling Unit). 

The TC communication between DHU and payload is summarized on the following figure, where N is the length of 

the TC in words of 16 bits (N ≤ 32) 

Message BC to RT 

Receive 

Command Data Word 1 Data Word 2 

... 

Data Word N-1 Data Word N 

Response RT to BC 

Status Word 

Figure 3.4-2: TC communication 




PL - 3.4.4 - 10 


Instrument unit delay constraint between two consecutive TCs shall be documented in its IDS. The delay shall 

be managed by ground using TTCs for instance. No specific delay other than time tagged execution is 

implemented on board. 

In particular, if 1553 TCs have to be processed in the same 250 ms cycle, they can be sent on 1553 in the same 

cycle without delay between each other. 

3.4.4.3.2 Shutdown 

Not used. 

PL - 3.4.4 - 11 

Deleted 




3.4.5 PAYLOAD TELEMETRY 


The observability function allows for monitoring of the satellite’s behaviour and groups all the mechanisms which 

collect data on the satellite, transmitting this information to the ground. From the on board software point of view, 

this starts with acquisitions up to delivery of telemetry packets to the telemetry down link hardware. 

PL - 3.4.5 - 1 

The average PLTM rate (1553) shall be lower than « mission dependent value » kbit/s. 

The rationale of this requirement is to size the autonomy of the satellite and the number of ground stations required. 

this value shall be determined by system mission analysis. 

PL - 3.4.5 - 2 

PLTM shall be filled by the 1553 bus. 

3.4.5.2 TM from discrete acquisitions 

3.4.5.2.1 General 

PL - 3.4.5 - 3 

Payload parameters acquisition is based on a cyclic concept. 

The frequency of the acquisition will be adaptable, depending on the payload. 

This frequency shall be defined as 1/32 Hz, 1/8 Hz or 1 Hz. 

Discrete acquisitions will be formatted as TM packets by the satellite on board software. 

These packets will be part of the platform housekeeping telemetry. 

3.4.5.2.2 Housekeeping telemetry 

Housekeeping Telemetry (HKTM) messages are devoted to advise ground of the on board events and to perform a 

cyclic monitoring of all the in use flight units including the payload. This telemetry contains cyclic TM packets 

recorded with pre-defined telemetry frequencies, and asynchronous TM packets generated « on events ». Payload 

HKTM is separated from platform HKTM by specific APID (Application Identifier) in each packet header. 

The downlink flow is divided into the following flows : 

The HKTM-P flow (permanent flow) : information of the current state of the satellite. This flow is a permanent 

flow, received in real time by the ground station when the satellite is in visibility and the emitter ON. 

The HKTM-R flow (registered flow) : the information contained in this flow is recorded in the on board mass 

memory and sent to ground upon request. It gives the « history » of the satellite for long term analysis and 

observability of some programmable « zooms » on particular events. 

HKTM packets will be adapted for each mission. 

Two HKTM packets contain payload power lines relay status and current and all discrete payload acquisitions as 

defined in chapter 3.4.1 : relay status, analog acquisitions, temperature acquisitions, DS16 data... 




3.4.5.3 TM from 1553 acquisitions 


PL - 3.4.5 - 4 

Scientific TM messages shall be formatted by the payload as CCSDS standard TM packets (packetisation 

layer). The TM messages shall have a maximum length of 512 words of 16 bits. 

These TM packets will be stored in the PROTEUS mass memory, before being downloaded to ground in visibility 

periods, on ground request. Payload telemetry is extracted from mass memory and transferred to TM downlink by 

OBSW, packet by packet. 

Two mass memory areas (PLTM1 and PLTM2) are saved for the payload telemetry; their size and mapping are 

configured at OBSW generation and are mission dependent. In case of overloading the storage capacity, the oldest 

TM are overwritten by the new ones. The maximum storage capacity for the PLTM is limited to 2 Gbits useful at the 

End Of Life. The PLTM granularity is equal to 16 Mbits. 

These two levels (PLTM1 and PLTM2) will be downloaded according to their priority (mission dependent) 

The storage mechanism of the payload telemetry (PLTM) packets depends on the maximum instantaneous data rate 

generated by the payload. 

During the Safe Hold Mode, the telemetry packet time reference is the current OBT. During all the other modes, the 

TM packets dating reference is UTC. 

In the PROTEUS standard, payload management is not authorized during Safe Hold Mode (RT are isolated). 

PL - 3.4.5 - 5 

According to their respective RT address and sub address, 1553 messages may be stored either in the PLTM1 

or PLTM2 . 

This choice is to be indicated within the ICD. 

PLTM storage are exclusive and cannot be modified in flight. 

If necessary, few packets could be stored in HKTM and this option shall be negotiated with ALCATEL SPACE or CNES. 

Nota : TM messages are written in Mass Memory according to the length indicated in the packet ; if this length is 

over 1024 bytes, the packet will be truncated to 1024 bytes and so the end of the message lost. 




3.4.5.3.1 Telemetry protocol (RT to BC transfer) 

3.4.5.3.1.1 RT to BC Functional requirements 

PL - 3.4.5 - 6 

The BC will regularly poll each Remote Terminal, by starting to send it a TRANSMIT_VECTOR_WORD 

mode_command. 

• The polling rate is set at 250 ms. 

• The RT shall accept the TRANSMIT_VECTOR_WORD MODE_COMMAND (mode_code 10000). 

• When needed, the RT shall request to transmit one or several messages by setting the 

SERVICE_REQUEST bit in the STATUS word. 

• The RT shall transmit the type of the messages it wants to transmit by setting relevant bits in the VECTOR 

word. 

It is the Payload Supplier responsability not to rise bits in the VECTOR WORD corresponding to undeclared 

subaddresses. This would induce in PLTM trains of bytes at 0000h to which the Payload Ground Segment may not be 

robust. 

Each bit in the vector word is defined as a subaddress ; that means that, in the proposed protocol, a given message 

corresponds to a given subaddress. 

PL - 3.4.5 - 7 

Each bit in this word corresponds to a type of message, that gives an implicit definition of the length of the 

message. 

• If the message length is less than or equal to 32 words, then a single_message scheme shall be used. 

• If the message length is more than 32 words and less than or equal to 512 words, then this message 

shall be acquired by consecutive 1553 messages of 32 words (or less for the last 1553 message). 

• Example : MSG = 512 words 

= 16x1553 messages of 32 words 

or MSG = 255 words 

= 7x1553 messages of 32 words and 1x1553 message of 31 words 

• SR bit shall be set to 1 when messages are ready for acquisition. 

At the end of the polling sequence, the BC will analyse the VECTOR word of each RT, compute the number of 

messages it is able to acquire during the allocated cycle, and program the BC to acquire the selected messages. 

PL - 3.4.5 - 8 

The RT shall accept an acquisition sequence close to the polling sequence. This means that messages shall 

be ready in the RT buffer before the RT rises its service request. 

Nota : Delay between vector word indicating the type of messages and «transmit data» command can be 0 s. 

Non selected messages will be acquired on the next cycle as far as service request bit is still raised. A given message 

(i.e. a subaddress content) is acquired during one single cycle (no splitting on two cycles). 




PL - 3.4.5 - 9 


When the acquisition frame is completed without error, the BC enters in an acknowledge sequence : 

• The RT shall accept a SYNCHRONIZE_WITH_DATA_WORD command (mode code 10001). 

• On the reception of the command, the RT shall reset VECTOR word bits. 

• The SR (Service Request) bit shall stay in the SET state until all messages have been acquired by the BC, 

or if new messages have to be read out. 

Example of SYNCHRONIZE_WITH_DATA_WORD command, according to preceding example of service 

request. 

MSB 0 

LSB 15 

Sub Address 10 still to be acquired 

Sub address 13 still to be acquired 

SYNCHRONIZE data WORD 

Sub address 1 : message 1 acquired if bit set to 0 

Sub address 5 : message 5 acquired if bit set to 0 

• The SYNCHRONIZE_WITH_DATA_WORD command will have the same structure as the RT VECTOR 

word. 

Bit convention : each bit of the SYNCHRONIZE_WITH_DATA_WORD command is set to zero by the BC when 

the corresponding message is fully acquired. 

Unused bit will stay to zero (subadress 0 and 15). 

The other bits shall be set to 1. 

• The delay between a SYNCHRONISE_WITH_DATA_WORD command and the next polling will be at least 

32 ms. 

Please note that for packets not declared in the IDS (thus unused on this RT), the bits of the data word will stay to 0, 

while for used subaddresses, the bits will be put to 1 if the message was not collected during the cycle or if nothing 

happened on this subaddress during the cycle (no request from the RT and no collection from the BC) and while 

messages collected during the cycle will have a bit set to 0. 

PL - 3.4.5 - 14 

The vector word shall be updated by the payload units according to 1553 protocol in less than 32 ms. 

3.4.5.3.1.2 RT to BC timing requirements 

The duration of the polling sequence is zero at minimum and is repeated every 250 ms. 




PL - 3.4.5 - 10 


Considering the proposed protocol, if some timing constraints are imposed by the RT (ex : elapsed time 

between the set-up of the service_request bit and the acquisition of the data), these constraints shall be 

reported within the RT IDS. 

Nota : Inter message gap (delay between the end of the last data word and the next transmit command, i.e. gap 

between messages) can be [8µs, 11µs] (1553 standard value). There are no other minimum or maximum timing 

constraints than the ones imposed by the standard protocol. 

The acquisition of the messages can occupy the bus without disturbing the BC host during about 140 ms (refer 

protocol limitations hereafter : maximal allocation for acquisition is 2560 words which represents 140 ms in a 250 

ms CPU cycle). 

About 32 ms are reserved to transmit Ground commands. 

3.4.5.3.1.3 Error handling on RT to BC transfer 

There will be no automatic repeat_procedure in case of protocol level 2 error (time-out,...) 

The RT will be put in STANDBY status by the OBSW (no longer polled regarding the 1553 communications, but 

observable via the discrete acquisitions), see PL-3.4.6-2. 

3.4.5.3.1.4 RT to BC Protocol limitations 

PL - 3.4.5 - 11 

The number of types of messages is limited to 14 max per RT (subaddresses 1 to 14). Two other 

subaddresses are reserved for PF use (broadcast...). 

PL - 3.4.5 - 12 

The full acquisition capacity (peak value) is 2560 words during a slot of 250 ms, independently of the 

sending RT (140 ms for acquisition/16 ms for one 512 word message). 

That means that the proposed protocol offers to sample : 

one RT updating 5 messages of 512 w every 250 ms, 

or 5 RT, each updating 1 message of 512 w every 250 ms, 

or any combination which does not overrun the above mentioned limitation. 

The maximum peak acquisition capacity is thus 2560 words for 250 ms (204 kbits/s, rate based on 20 bits words : 

16 useful bits + 1 parity bit + 3 synchronisation bits, without taking into account the transmit command and the 

status word). 

If RTs messages total for a slot of 250 ms (one cycle) has a length over 2560 words, the OBSW takes into account 

every RT message in decreasing priorities order until a total of 2560 words for the considered cycle. The other ones 

not selected in this cycle are taken into account in the next cycle. 

Nota : «2560 words selection» principle 

After polling, the BC chooses packets: it begins its selection with packets having priority until 2560 words limit. If a 

packet is too long to be chosen, the OBSW goes on the selection with packets having less priority to reach the 2560 

words limit without exceeding. The «no selected» packet will be taken into account in the next selection processus. 

PL - 3.4.5 - 13 

It is the responsibility of the Payload Supplier to ensure that the rate is not overrun. 




The TM communication between payload and DHU is summarized on the following figure. The OBSW is in charge of 

building the interrogation frame related to the complete instrument packet acquisition. 

This will be done with as much interrogations commands as needed: to acquire a packet with a maximum length of 

512 words (header included), 16 interrogations will be sent to the RT. 

It is not requested that packet length is a multiple of 32 words; OBSW manages the length of the 1553 messages 

according to packet lengths. 

The header of TM packets is 8 words of 16 bits long (only the first 3 words are defined by the CCSDS standard, 

datation is added). 

Message 1 : BC to RT 

Transmit 

Command 

Response 1 : RT to BC 

Status 

Word 

Message N : BC to RT (2 ≤ N ≤ 16) 

Transmit 

Command 

Response N : RT to BC 

Status Word 

Header 

Word 1 

Data 

Word 1 

... Header Data ... 

Word 8 Word 1 

Header word 1: 3 bits forced to 100 version number (CCSDS standard) 

1 bit forced to 0 type indicator 

1 bit forced to 1 secondary header presence 

11 bits application identifier specific to the payload 

Data 

Word 2 

Data 

Word 31 



... 

Data 

Word 24 

Header word 2: 2 bits forced to 11 grouping flag 

14 bits sequence count 

these bits may be set at instruments convenience 

Header word 3: 16 bits 

Header word 4 to 8: 

packet length, number of bytes of the packet data zone 

minus one 

80 bits date coded on 10 bytes according to section 3.4.7 

bits 13 and 14 of the two week number bytes may be set 

at instruments convenience 

Figure 3.4-3 : Scientific Telemetry Exchange 

Data 

Word 32


3.4.5.4 Deleted 





3.4.6 PAYLOAD SURVEILLANCE 

PL - 3.4.6 - 1 a 

The payload supplier shall discuss its monitoring and surveillance need with the Satellite Contractor.. 

The platform offers nominally payload surveillance at 1/32 Hz and 1/8 Hz. It may also provide surveillance 

at 1 Hz. 

However the surveillance shall be limited to current, temperature and analogic parameters including voltage 

(not including thermal control lines surveillannces provided by the platform if any) 

The number of these surveillances shall be lower than 16. 

Surveillance could be also offered with logical status and serial lines affected to the instrument (mask on 16 bits). 

Standard design includes also a surveillance (one alarm temperature) on each thermal control line dedicated to 

payload (if any). Refer to § 3.2.2 option 2. 

In addition, 9 thermistors in spare leading to SHM could be used by the Payload. 

3.4.6.1 Failure Detection Isolation and Recovery 

The failure detection function is in charge of the satellite monitoring in order to keep it in a safe state. 

The operational monitoring is performed by the OBSW. In case of a software malfunction, an hardware watchdog 

triggers and an hardware automaton are able to reconfigure the satellite units. 

The isolation function satisfies the following requirements : 

On board automated tasks are in charge of the decision of switching the satellite to Safe Hold Mode. 

In baseline, no isolation, but only deactivation of a failed unit is done on board. 

The recovery function is split between on board automatism and ground interventions : 

Safe functions are recovered by on board automated tasks. 

Unit recovery is performed by ground. 

Commandability should be as simple as possible. 

These principles imply the following rules : 

Every time a critical platform failure is detected on board (including the payload thermal control surveillance), 

the satellite is switched to Safe Hold Mode and the mission is interrupted. All the payload units are switched 

OFF (except the 2 lines «8» and «16» which may be ON) as the other in use platform units and then, the 

satellite is switched to Safe Hold Mode and the payload units are in ISOLATED status. 

Every time a critical platform failure is detected on board (including the payload thermal control surveillance), 

the satellite is switched to Safe Hold Mode and the mission is interrupted. All the payload units are switched 

OFF (except the 2 lines «8» and «16» which may be ON) as the other in use platform units and then, the 

satellite is switched to Safe Hold Mode and the payload units are in ISOLATED status. 

Every time a payload failure is detected on board, the concerned instruments or equipment are switched to 

their PASSIVE status and are switched OFF, but the satellite remains in normal mode. 

Every time a 1553 dialogue anomaly is detected, the corresponding instrument is autonomously turned in 

STAND BY status. 

In flight, functions build telemetry messages with computed or acquired parameters so that ground diagnostic 

is made possible. 




For each instrument, the main parameters that need being checked can be different. It is recommended to limit onboard 

automatic survey to the ones needing a quick intervention loop. Other parameters are monitored on the 

ground level. 

The parameters needing monitoring on ground request (FDIR activation) instrument are usually the following ones: 

the power line current (minimum and maximum limits) 

the temperature monitoring line affected to this instrument 

the analog parameters affected to this instrument (including voltage) 

PL - 3.4.6 - 2 

Finally, the payload shall withstand without any damage all consequences of a spacecraft FDIR action (power 

cut-off and/or MIL-STD-1553B dialog interruption) that is to say : 

• rough power cut, because SHM transition (all instruments are concerned except for the 2 specific lines 

described in section 3.5.3), or because an instrument failure (only this instrument is concerned even if it 

is connected to one of the 2 specific power lines). The 1553 dialogue is also interrupted. 

• turning in STAND BY status every time a 1553 dialogue anomaly is detected. 

• other actions depending on the mission need and payload specific surveillance. 

PL - 3.4.6 - 3 

If necessary, recovery action after preceding actions shall be indicated by the Payload Supplier in the Payload 

User’s Manual. 

PL - 3.4.6 - 4 

deleted 

3.4.6.2 Payload switch-off in case of SHM 

3.4.6.2.1 On PF failure detected by H/W 

PL - 3.4.6 - 5 

In case of system monitoring by hardware (DHU problem, TCD RM alarm, TCD warm reset) leading to SHM 

(PL anomaly are not included here and are considered in section 3.4.6.3), the payload will be isolated and 

its power lines will be switched OFF by groups as shown in table 3.4-1. The PL shall withstand this switch 

OFF sequence. 

Sequence duration Payload lines group # 

26 ms P1 (1 to 4) 

3 ms 

26 ms P2 (5 to 7) 

3 ms 

26 ms P3 (9 to 12) 

3 ms 

26 ms P2 (13 to 15) 

Table 3.4-1: Payload switch-off sequence by reconfiguration module (see figure 3.5-7 for lines group 

definition) 




3.4.6.2.2 On PF failure detected by S/W (including generic PL CTA failure) 

PL - 3.4.6 - 6 a 

In case of system monitoring by software leading to SHM (PL anomaly is not included here 

and is considered in section 3.4.6.3), the payload will be switched off following a generic 

sequence defined hereafter. The payload shall withstand this switch OFF sequence. 

• The faulty platform surveillance recovery action is inhibited. 

• 500 ms delay * 

• DHU internal group relays are switched OFF by Reconfiguration Module (refer to PL-3.4.6-5) 

• new SHM 

* This parameter belonging to the SDB could be re-configurated. 




3.4.6.3 Payload switch-off in case of payload anomaly 

PL - 3.4.6 – 7 b 

Payload surveillance (description, max. filter value, monitor frequency) leading to FDIR mechanism will be 

defined by the payload supplier as defined in PL-3.4.6-1. 

It will lead to the following platform actions under platform software control: 

• firstly, the power relay of the faulty payload instrument will be opened (the operation will be done before 

the «n+1» monitoring cycle, «n» being the filter value in case of analogic data) 

• then, the payload instrument is set to a PASSIVE status. 

The payload shall withstand this switch OFF sequence. 

In case of 1553 dialog failure the payload instrument is only set to a STANDBY status. 

The following figure illustrates the previous requirement. 

Figure 3.4-7: Payload switch-off timing in case of payload anomaly («n» being the filter value) 




3.4.7 TIME AND SYNCHRONISATION DISTRIBUTION 

The PROTEUS Platform GPS is in charge of the delivery of a precise synchronisation pulse each second, so called PPS 

(pulse per second ; the PPS is distributed to the payload via the DHU by hardware design). 

PL - 3.4.7 - 1 

Eight PPS signals (pulse per second) are available for the payload (each unit will be shared, that is to say 

accessible by the nominal PM and by the redundant PM). 

These pps signals are only available when the GPS is ON ; guaranteed when the platform is in nominal mode (CC 

mode), in normal operating mode. 

In reduced mode (CC mode), the pps performance is damaged : pps frequency depends on the GPS drift and pps 

datation depends on the OBT drift. 

For Nom to reduced transition, a biases of Nx50 ms (with N : programmable) can appear on the pps datation. 

PL - 3.4.7 - 2 a 

The payload shall be compatible with the fact that these signals are absolute references (generated by an on 

board GPS receiver giving the precise date associated with the coming PPS) in UTC and that they are phased 

on entire seconds, coded on 10 bytes and have an accuracy better than 5 µs. Moreover, the date will be 

delivered to the payload by the SW of the DHU via the 1553B data bus by a broadcast command. 

PL - 3.4.7 - 4 a 

The payload shall be compatible with the fact that the datation format in the second header of the platform 

header will be the following, on 10 bytes: 

• week number : integer, 2 bytes (bits 4 to 15) ; LSB is 1 week. It is to be noted that the first week (6 to 12 

January 1980 is numbered zero (0), 

• seconds in the week : long integer, 4 bytes. LSB is 1 second. 

• second fraction within the second : long integer, 4 bytes. LSB is 2 -32 second. 

The week number is encoded on the least significant 12 bits of the first word of the time bulletin (bits 4 to 15). The 

time source field is encoded on the most significant bit of the first word. Bit 1, 2and 3 are unused; bit 0 indicates 

UTC (if set at 0) or On-Board Time (if set at 1). 

The reference date (0H) is: 6 january 1980, 0h00 if UTC, or indicates the beginning of the satellite Safe Hold Mode 

if On Board Time is used. 

Bit 0 represents datation source, either UTC from GPS if GPS is functional (value 0) or derived from OBT if GPS has 

not been yet started or if a GPS constellation problem prevents from using it presently (value 1). Please note that first 

case (value 0) corresponds to satellite Command Control NOMINAL mode, while second case (value 1) corresponds 

to Command Control SAFE or REDUCED modes. In this latter case (REDUCED), last available GPS date or a ground 

issued date bias allow to ensure datation continuity, however with reduced performance, since on board oscillator 

does not match the GPS date accuracy on the long term. However, such a situation has to be considered only as a 

transient. 

PL - 3.4.7 - 3 a 

PPS and polling sequence are not synchronised. Consequently, the payload shall be compatible with the 

delivery of a broadcast message containing the date of each pulse delivered 825 ms to 250 ms before PPS 

delivery. 

The broadcast command is the last message sent within a 1553 cycle. 




Figure 3.4-7: Time bulletin versus PPS signal 

The PPS electrical interfaces are described in the chapter 3.5.6.3. 

Each HKTM packet contains the date of packet generation using the same UTC reference in the first packet data 

words. 




3.4.8 PAYLOAD SPECIFIC SOFTWARE INSIDE DHU 

As an option, a payload specific software can be loaded inside the DHU. The used language is ADA. Software 

interfaces will be discussed with ALCATEL SPACE and CNES (the available volume and the percentage of the CPU 

load dedicated to the payload will be negotiated). 




3.5 ELECTRICAL INTERFACE REQUIREMENTS 

3.5.1 GENERAL SYSTEM CONFIGURATION 

Power supply and command control is performed by the data handling unit (DHU). 

The DHU is based on the cold redundancy concept of two processor modules (PM). 

Power lines and command/control lines (including pyro lines) are organized in independent interfaces modules 

accessible by the two PM. These interface modules have a high reliability. 

Power lines are protected by fuses within the DHU. 

Switching of power lines is performed inside the DHU. 

Payload pyro lines are protected by three safety barriers inside the DHU. 

Command and control is implemented either via busses compliant with the MIL-STD-1553B standard and/or via 

direct commands. 

A payload unit can be connected to the DHU, as shown on the Figure 3.5-1, Figure 3.5-2 and Figure 3.5-3 : 

.. 

DHU 

PM A 

PM B 

IF A 

IF B 

Power A 

Power B 

inst. side A 

inst. side B 

DHU internal Bus 

Figure 3.5-1 : System configuration with redundant units and cross-strapping 



TM/TC 

MIL STD 1553 

Power lines

.. 


DHU 


DHU 

PM A 

PM B 

PM A 

PM B 


IF A 

IF B 

Power A 

Power B 

instrument 

Figure 3.5-2 : System configuration with single unit internally redundant 

IF A 

IF B 

Power A 

Power B 

instrument 


Figure 3.5-3 : System configuration with no redundant unit 



TM/TC 

MIL STD 1553 

Power lines 

TM/TC 

MIL STD 1553 

Power lines


3.5.2 PIN ALLOCATION 

PL - 3.5.2 - 1 

The Payload shall comply with the following connectors brackets description (H01, H02 and H03) and the 

pin allocation given in appendix B. 

Note: J designation is given for fixed connector, P for mobile connector. 

The lines not used for the Payload will not be physically wired on Payload side. 

3.5.2.1 Power bracket 

The connector description is given through Figure 3.5-4 and Table 3.5-1. 

Detailed pin allocation is given in appendix B. 

Connector 

Code 

Payload Interface Power Connector Bracket 

Platform DHU Payload Nominal Thermal control heaters 

Platform DHU Payload Nominal Power 

Platform DHU 

Platform DHU 

Platform DHU 

Platform DHU 

Payload Nominal Pyros lines 

STR1 Star Tracker 1 Power 

STR2 Star Tracker 2 Power 

Payload Redundant Pyros lines 

Platform DHU Payload Redundant Power 

Platform DHU Payload Redundant Thermal control heaters 

Number 

of pins 

H01 

Figure 3.5-4 : H01 Connector bracket 

Sex 

(M/F) Description Connector ref 

Payload Power Wiring connectors (to H01) 

P01 25 M Nominal Thermal control heaters DBM-25P 

P02 37 M Nominal Power DCM-37P 

P03 37 M Nominal Pyros lines DCM-37P 

P04 9 M Star Tracker 1 Power DEM-9P 

P05 9 M Star Tracker 2 Power DEM-9P 

P06 37 M Redundant Pyros lines DCM-37P 

P07 37 M Redundant Power DCM-37P 

P08 25 M Redundant Thermal control heaters DBM-25P 

Table 3.5-1 : H01 connector bracket description 




3.5.2.2 Acquisition and command interface brackets 

The connector description is given through Figures 3.5-5 and 3.5-6 and also through Table 3.5-2 and 3.5-3. 

Detailed pin allocation is given in appendix B 

Nominal Payload Interface TM/TC 

k 

Payload


Redundant Payload Interface TM/TC Connector 

B k t 

Payload


3.5.3 POWER LINES 

3.5.3.1 Available lines 

PL - 3.5.3 - 1 b 

Payload power interface is given by the DHU which typically provides 16 switchable unregulated power lines 

(5A max) on the unregulated bus. This distribution permits 16 lines simultaneously or to have a cold 

redundancy concept at Payload /platform level (cf. Figure 3.5-7). 

One instrument is powered by one power line. 

Power 

Bus 

Figure 3.5-7 : DHU I/O channel layout 

These power lines are distributed as the Command and Control resources (§3.4.1), that is to say in three 

independent groups (P1, P2, P3) inside the DHU : 

Lines number 

P1 1, 2, 3, 4, 8 

P2 5, 6, 7, 13, 14, 15 

P3 9, 10, 11, 12, 16 

Remark: The 3 independent groups (P1, P2 & P3) indicated on the Table above correspond with respectively the 3 

independent cards (SiOP 3, SiOP 4 and SiOP 5) implemented inside the DHU. 

This distribution might be considered for the redundancy philosophy. 

There are 2 independent levels of relay to command these payload lines for safety reasons. One level is commanded 

by software with standard TC, and the other level is directly commanded by hardware TC from ground or 

reconfiguration module except for lines 8 and 16 for which both relais are commanded by software, but through 

independent hardware electronics. The ON/OFF status of each line is accessible to software and telemetry. The 

current consumption on each line is accessible to software and telemetry. 

Only current (not voltage) is monitored and activates alarm on max threshold. Relay status are also monitored. The 

16 power lines are concerned whatever they are used or not. Standard monitoring frequency is 1/8 Hz. The primary 

voltage is also monitored at battery level (not at each line level) (note that as far as the resistance of line is known, 

the voltage of each line is also known). These monitoring are performed by the platform and information are 

included in a devoted HKTM packet. 




PL - 3.5.3 - 2 


During launch phase, the power lines 8 and 16 can be configured before launch, to supply power to part of 

the payload if needed (TBC depending on launch phase). But in this case they are not controlled by software 

before separation due to the OFF status of the data handling unit processors. 

These two specific power lines remain in the same status in case of Safe Hold Mode transition (after on 

board failure detection). So a payload , powered by these specific lines, which is ON before a transition 

towards SHM will be maintained ON during and after reconfiguration. Global payload consumption of these 

two power lines has to be limited to 30 W. 

As an option (mission specific), output lines might be merged and sized to get a power supply line of more than 5 A. 

Nota: However this configuration is not recommanded during launch. Compatibility with the launch vehicle shall be 

analysed in this case. 

PROTEUS platform offers several solutions to fit with the payload electrical requirements. The User is cordially invited 

to contact ALCATEL SPACE and CNES in order to optimise the electrical distribution at payload/platform level 

considering power and energy budgets for its mission. 

3.5.3.2 Payload power consumption 

PL - 3.5.3 - 3 a 

Power budget shall be established at a voltage of 28 V. 

PL - 3.5.3 - 4 a 

Power consumption and dissipation shall also be provided at the BNR values voltage 23 V (relevant lines), 28 

V and 37 V. 

PL - 3.5.3 - 5 

The nominal average power consumption, the variation and the dispersion values and peak power demand 

shall be updated periodically and are taken into account to establish satellite system budgets. 

An overstepping of the maximum power consumption toward the allocated power consumption will induce a formal 

power change notice 

3.5.3.3 Power interface characteristics 

The reference point for the following characteristics is at the H01 bracket. 

3.5.3.3.1 Input voltage 

The satellite will provide power lines for the payload with a voltage range of 23 V to 37 V DC unregulated power. 




PL - 3.5.3 - 6 a 


The payload shall operate within the two voltage ranges : 

• a nominal voltage range including one battery pack lost for Payload design & performance optimization 

(for all units never ON on SHM phases), 

• a degraded voltage range for which the Payload shall operate without any dysfunction and out of 

specification performances (for units wich must be ON on SHM phase via the power lines 8 and 16) 

The corresponding voltage ranges are the following : 

• nominal voltage : [27.5 V --> 37 V] 

• degraded voltage range : [23 V --> 37 V]. 

3.5.3.3.2 Source impedance 

PL - 3.5.3 - 7 

deleted 

The distributed users line output maximum impedance is the primary bus impedance degraded by 3 dB for 800 Hz< 

f < 60 kHz (computation gives 447 mOhm for 4 kHz < f < 6 kHz and 355 mOhm for 6.1 kHz < f < 60 kHz). 

1,00E+01 

1,00E+00 

1,00E-01 

1,00E-02 

Z (Ohm) 

1,00E+00 1,00E+01 1,00E+02 1,00E+03 1,00E+04 1,00E+05 1,00E+ 

Figure 3.5-8 : DHU output impedance 




3.5.3.3.3 Transient 

PL - 3.5.3 - 8 

The maximum variation of the nominal DC voltage of the primary bus (at DHU input) and of the distributed 

switched lines (at DHU output) is 55 V DC during 0.5 ms and 42 V DC during 5 ms. The operational 

transients voltage variation does not exceed ±0.5 V DC for a duration less than 5 ms. Steady state will be 

reached after this duration. 

3.5.3.3.4 Input voltage ripple 

PL - 3.5.3 - 9 

The input voltage ripple and spike, including the effects of all loads, will not exceed 1.0 V peak-to-peak in 

the range 50 Hz to 10 MHz. 

3.5.3.3.5 Power lines protection 

PL - 3.5.3 - 10 

Deleted 

Power lines protection will be accomplished by the use of fuses within the DHU (Data Handling Unit) for the + 

primary power lines. 

PL - 3.5.3 - 11 a 

The Payload design shall be compatible with 10 A fuses on each power line (corresponding to 5 A maximum 

current with a rating factor of 2). 

The equipment shall not be degraded by the transients on the power bus due to fuse blowing as indicated in 

Figure 3.5-29. 



with 


PL - 3.5.3 - 12 

Figure 3.5-29: Transients of the power bus to fuse blowing 

The use of fuses inside payload is forbidden. If necessary, a specific protection (example : currrent limiter) 

shall be implemented inside the payload to limit effects of short circuit and avoid transients greater than 

specified. 

PL - 3.5.3 - 13 

It is recommended that no primary power relay be implemented inside the payload except in case of 

merged power lines. 

3.5.3.3.6 Input current 

PL - 3.5.3 - 14 

The maximum allowed current consumption on payload power line is 5 A DC. 

3.5.3.3.7 Converter input impedance 

PL - 3.5.3 - 15 

Power converters within the payload shall have an input impedance electrically adapted to the distributed 

switched lines output impedance (see 3.5.3.3.2). 




3.5.3.3.8 Undervoltage 

PL - 3.5.3 - 16 a 

The payload and associated components shall not be damaged by any input voltage on the range [0 V --> 

Minimum Voltage value as defined in PL-3.5.3-6]. 

3.5.3.3.9 Overvoltage 

PL - 3.5.3 - 17 

The payload shall withstand the power transients defined in section 3.5.3.3.3. 

3.5.3.3.10 Reverse polarity 

PL - 3.5.3 - 18 

Payload shall not be damaged after power connection with reverse polarity. 

3.5.3.3.11 Short circuit protection 

PL - 3.5.3 - 19 

Deleted 

PL - 3.5.3 - 20 

Payload connected to nominal and redundant power lines shall have an internal isolation of the lines by « 

OR » diodes or equivalent devices. 




3.5.4 PYROTECHNIC LINES 

Pyrotechnic lines are organised in independent interface modules accessible by the two processor modules of the 

DHU. 

PL - 3.5.4 - 1 

16 pyrotechnic lines (8 nominal and 8 redundant) are dedicated to the payload. 

PL - 3.5.4 - 2 

For the payload pyrotechnic lines, the pyrotechnic function is performed after three independent barriers: 

barrier 1, barrier 2 & selection relays and one current limiter (ensuring the same function as a firing relay) 

according to its mission (cf. Figure 3.5-9) 

The barrier 1: electromechanical relay. This relay, common to 8 platform and 8 payload lines, is inhibited by a 

separation strap coming from the umbilical. This latching relay is commanded by the software from the processor 

module. 

The current limiter: this current limiter transistor is common to 8 platform and 8 payload lines, is commanded by 

software from the processor module (pulse duration 26 ms). The current limitation is rated to 5 A. 

The barrier 2: electromechanical relay. This relay is commanded by hardware telecommand sent by ground 

command. 

The selection relay : electromechanical relay. This relay is commanded by software command from the processor 

module. 

Figure 3.5-9: Electrical inhibit implementation (only the main branch is illustrated) 

PL - 3.5.4 - 3 

The electroexplosive initiator shall be NSI or equivalent, type 1 A / 1 W / 5 mn NO FIRE. 

Electrical characteristics are: 

No fire current: 1.0 A during 5 minutes until 150 °C. 

Bridgewire resistance: 0.95 to 1.15 Ω. 




Checkout current: 0.01 A maximum for 1 minute max. 

Electrostatic sensitivity: not degraded by 25 kV from 500 pF capacitor and 5 kΩ series resistor. 

Insulation resistance: 2 MΩ mini at 250 V (DC) between initiator shorted pins and case,15 s 

minimum. 

Recommended firing current: 5 A (DC), 10 ms from -162 °C to +150 °C. 

Minimum firing current: 3.5 A, 10 ms from -54 °C to +150 °C. 

Leakage current: 

PL - 3.5.4 - 4 

less than 0.050 A at 28 V DC after firing. 

The payload pyro lines commanding hazardous devices shall be protected by a safe and arm plug in the 

payload harness located on +Z payload panel, as close as possible to the –X payload interface plane and 

accessible at any time during satellite integration. 

PL - 3.5.4 - 5 

The payload pyro lines safe plug and arm plug connection procedure shall be given by the payload supplier 

and approved by ALCATEL SPACE 

This protection may be achieved by installing a Safe plug in arm plug receptacle, or by intrinsic design of the firing 

circuits. 

The activation of the pyro lines and the control of the overall sequence may be done using time-tagged TC from 

ground or using a specific application in on-board software. 

PL - 3.5.4 -12 

Arm and Safe Plugs or caps shall be designed to be positively identifiable by color, shape and name. The 

natural body color of the Arm plug is required. The safe plug or cap should be green and shall have a red 

remove- before –flight streamer attached. They shall be marked Arm and Safe respectively. 

PL - 3.5.4 -13 

The design of the device and the firing circuit shall ensure easy access for plug installation and removal 

during assembly and checkout in all prelaunch and post- launch processing facilities. 

PL - 3.5.4 -14 

Monitor and control circuits shall not be routed through Safe Plugs. 

PL - 3.5.4 -15 

For each pyro line the overall resistance from H01 to the pyro shall be less than 7ohms. 

PL - 3.5.4 -16 

Electroexplosive devices shall be protected from electrostatic hazards by the placement of resistors from line 

to line and from line to ground (structure) 

The placement of line to structure static bleed resistances is not considered to violate the single point ground 

requirements of this standard as long as the parallel combination of these resistors are 10 K omhs or more. 

PL - 3.5.4 -10 

The Electro Explosive Device Extension Subsystem shall be designed to limit the power produced at each EED 

by the electromagnetic environment acting on the subsystem to a level at least 20dB below the maximum 

pin-to-pin DC no fire power of the EED. 




PL - 3.5.4 -11 


The Electro Explosive Device Extension Subsystem shall be designed to limit the power produced at each 

devise in the firing circuit that complete any portion of the firing circuit to a level at least 6dB below the 

minimum activation power for each of the safety devices 

PL - 3.5.4 -17 

There shall be no aperture in any container which houses elements of the firing circuit. 

PL - 3.5.4 -18 

Application of operational voltage to the monitor circuit shall not compromise the safety of the firing circuit 

nor cause the electroexplosive subsystem to be armed. 

PL - 3.5.4 -19 

The monitoring shall be dual fault tolerant against EED firing. 

PL - 3.5.4 -20 

Monitor circuits and test equipment that applies current to the bridgewire shall be designed to limit the open 

circuit output voltage to one volt. 

PL - 3.5.4 -21 

Monitoring currents shall be limited to one tenth of the no-fire current level of the EED or 50 milliAmps 

whichever is less. 

PL - 3.5.4 -22 

Electromechanical and solid-state switches and relays shall be capable of delivering the maximum firing 

circuit current for a time interval equal to ten times the duration of the intended firing pulse. 

PL - 3.5.4 -23 

The return lines grounding is provided by the DHU. 




3.5.5 THERMAL LINES 

3.5.5.1 Active thermal control 

The active thermal control algorithm, the number and the power of these lines are described in section 3.2.2. 

In addition with these data, some specific electrical aspects of these thermal lines are given in section 3.5.6.2.2. 

3.5.5.2 Thermal monitoring 

See section 3.5.6.2.2. 




3.5.6 COMMAND AND CONTROL LINES 

PL - 3.5.6 -28 

The payload when not powered shall not be degraded when receiving platform electrical signals as defined 

in the following paragraphs (characteristics defined in the DHU output tables). 

The electrical characteristics of payload at the bracket interfaces when not powered shall be within the user 

interface fault voltage signals as defined in the following paragraphs. 

3.5.6.1 Commands 

3.5.6.1.1 EED 

See section 3.5.4. 




3.5.6.1.2 Relay command (High Level Command and RF High Level Command) 

This type of command is intended to drive high power loads, such as relays. 

PL - 3.5.6 - 1 

It is a single positive voltage pulse. 

Successive high level commands, RF high level commands, or low level commands (see 3.5.6.1.3) shall not 

be issued before the end of the previous command pulse, whatever is this pulse (HLC, RF HLC, or LLC). 

That is to say no HLC, RF HLC or LLC within 28 ms following a HLC or LLC order ; and no HLC, RF HLC or LLC 

within 270 to 520 ms following a RF HLC order (depending on this RF HLC length). 

PL - 3.5.6 - 2 

DHU output 

The payload shall respond to relay command inputs having the following characteristics : 

Type of source Single ended positive pulse 

Voltage level, active 

22 V 162 Ω) 

Voltage level, passive 

0 V 

Pulse length (tp) 

26 ms ± 2 ms 

for HLC 

adjustable from 250 ms 

to 500 ms ± 20 ms 

for RF HLC 

Rise time (tr, 10% to 90%) 50 µs < tr < 500 µs (when loaded with 

Fall time (tf, 10% to 90%) 50 µs < tf < 500 µs 270 Ω || 0.6 nF) 

Driving current capability ≥ 180 mA maximum current for PL relay 

Sinking current 

≤ 50 µA 

Short circuit current 

≤ 400 mA protection against short circuit 

Output capacitance < 50 pF 

Fault voltage (emission) 

0 V 

(tolerance) 

Inductive spike 

short circuit to ground 

clamped by user diode 

Table 3.5-4 : Electrical characteristics of the DHU HLC output interface 




tr 

tf 

USER input 

PL - 3.5.6 - 3 

90% 

50% 

10% 

tp 

90% 

50% 

10% 

Figure 3.5-10 : Signal wave shape for the HLC pulses 

The HLC or RF HLC user input interface shall be according to the following characteristics : 

Input voltage 21.5 V 

Load impedance > 162 Ω 

Fault voltage 

PL - 3.5.6 -18 

(tolerance) 

(emission) 

0 V 


Table 3.5-5 : USER High Level Command input interface 

These commands shall be documented in these units electrical ICD and control/command IDS. 




3.5.6.1.3 Low Level Command 

This type of command is primarily intended to drive small loads. 

PL - 3.5.6 - 4 

The low level commands are of SBDL type. 

The Standard Balanced Digital Link (SBDL) is a fully differential interface, with a "true line" and a "complementary 

line" (see figure below). The status of the signal is defined as high when the true line has a positive voltage "1" level 

with reference to the ground and the complementary line has a "0" level with reference to the ground. The signal is 

defined as low when the true line is at "0" and the complementary line is at "1". 

DRIVER 

+ 

- 

DHU output 

Figure 3.5-11 :Electrical interface for the SBDL 



RECEIVER 

Circuit type Complementary driver 

High level (true line, ref. to ground) 4.0 V to 5.5 V (with 10 kΩ diff. Load || 1.2 nF and 

100 kΩ common load) 

Low level (true line ref. to ground) 0 V to 0.5 V 

Rise/fall times, tr/tf (10% to 90%) 0.2 µs to 1.0 µs 

(when loaded with 1.2 nF || 10 kΩ) 

Skew between inverted and < tr/4, and (with 10 kΩ diff. load, and 

non-inverted outputs 

< 100 ns 100 kΩ common load) 

Diff. output impedance 229 Ω < Z < 305 Ω for f ≥ 300 kHz, 

229 Ω < Z < 310 Ω for f ≥ 52 kHz, 

229 Ω < Z < 474 Ω for f ≥ 0 Hz 

Short circuit current < 100 mA 


-12 V 

(tolerance) 

-17 V 

Table 3.5-6 : Electrical characteristics of the SBDL complementary driver 

+ 

- 

+


User input 


Circuit type Differential receiver 

High level 

> +1 V (differentially) 

Low level 

Receiver input impedance 

< -1 V (differentially) 

differential 

120 Ω in series with 100 pF || > 10 kΩ || < 150 pF * 

single input to ground > 16 kΩ || < 150 pF 

Hysteresis ≥ 1 V 

Fault voltage (emission) -16 V 1.5 kΩ) 

(tolerance) 

* see figure 3.5-11a 

-13 V 

PL - 3.5.6 - 5 

Table 3.5-7 : Electrical characteristics of the SBDL complementary receiver 

120 Ω 

100 pF 

> 10 kΩ 

< 150 pF 

Figure 3.5-11a: Receiver input impedance (differential) 

The pulse duration shall be 26 ms ± 2 ms (the pulse is a positive pulse). 

Successive High Level Commands, RF High Level Commands (see 3.5.6.1.2), or Low Level Commands shall not be 

issued before the end of the previous command pulse, whatever is the pulse (HLC, RF HLC, or LLC).That is to say no 

HLC, RF HLC or LLC within 28 ms following a HLC or LLC order ; and no HLC, RF HLC or LLC within 270 to 520 ms 

following a RF HLC order (depending on this RF HLC length). 

Pulse characteristics: 

- quiescient : negative level (-5 V) 

- active : positive level (+5 V) 

Here is the [true line-complementary line] voltage shape: 

+5 V 

-5 V 

Figure 3.5-12 : Signal wave shape for the LLC pulses 



tp


PL - 3.5.6 -19 


These commands shall be documented in the units electrical ICD and control/command IDS. 




3.5.6.1.4 16 or 8 bits serial command 

PL - 3.5.6 - 6 

The serial commands (CS8/16) signals shall comply with the interface requirements of the SBDL differential 

driver/receiver interface, specified hereafter. 

The Standard Balanced Digital Link (SBDL) is a fully differential interface, with a « true line » and a « complementary 

line » (see figure below). The status of the signal is defined as high when the true line has a positive voltage « 1 » 

level with reference to the ground and the complementary line has a « 0 » level with reference to the ground. The 

signal is defined as low when the true line is at « 0 » and the complementary line is at « 1 ». 

.. 

DRIVER 

+ 

- 

DHU output 

Figure 3.5-13 : Electrical interface for the SBDL 

DRIVER RECEIVER 

Data DHU USER 

Clock DHU USER 

Enable DHU USER 

Table 3.5-8a: DRIVER and RECEIVER values for Data, Clock and enable 



RECEIVER 











229 Ω < Z < 310 Ω for f ≥ 52 kHz, 

229 Ω < Z < 474 Ω for f ≥ 0 Hz 




(tolerance) 


Table 3.5-8 :Electrical characteristics of the SBDL complementary driver 

+ 

- 

+


User input 



High level 


Low level 



differential 





(tolerance) 


-13 V 


PL - 3.5.6 - 7 

The CS command timing is according to the figures below. The values specified are valid at the DHU output 

when it is loaded with 1.2 nF || 10 kΩ. 

Address 

(sample) 

ML Clock 

ML Data 

t2 

t1 

t4 t5 

t6 

B0 B1 B2 B3 B4 B5 B6 B7 B8 B9 B11 B13 B15 

B10 B12 B14 



t3 

t2+t4 

t7 t8 t9 

t8 t9 

Figure 3.5-14 : Serial Command timing (B0 is MSB) 

t1


t1 64 t ± t 

t2 4 t ± t 

t3 124 t ± t 

t4 28 t ± t 

t5 4 t ± 0.1 t 

t6 2 t ± 0.25 t 

t7 < 4 t 

t8 2 t ± 0.25 t 

t9 > 1.5 t 

Rise time 0.2 µs < tr< 1.0 µs 

Fall time 0.2 µs < tf< 1.0 µs 

where t = 2 -20 s ≈ 0.95 µs 

Table 3.5-10 : Characteristic Times values 

Rise and fall times are valid for all three signal types : address, clock, and data. 

Output data from the DHU is changed on the rising edge of the clock. Sampling of data shall be performed by the 

User receiver on the falling edge of the clock for maximum timing margins. 




3.5.6.2 Telemetry 


3.5.6.2.1 Analog telemetry 

DHU input 

PL - 3.5.6 -20 

The instrument active analog channels characteristics shall be compatible with the DHU active analog input 

interface described hereafter: 

Type of receiver 

Input voltage range 

Resolution 

Absolute accuracy 

Shielding 

Input Impedance 

differential 

differential 

single input to ground 


Common Mode Rejection Ratio 

(CMRR) for voltage - 2 V 10 MΩ (at power ON) 

≥ 1 kΩ || ≤ 1 µF (at power OFF) 

60 dB up to 10 kHz 

falling 20 dB/dec up to 1 MHz 

Leakage current < 1 µA (at power ON) 

< 0.5 mA (at power OFF) 



(tolerance) 

-14 V 

Table 3.5-11 : Electrical characteristics of the DHU AN input interface 

Note: As only the unipolar mode is to be used by the PL, the input voltage is positive. Nevertheless, the general 

bipolar description calls for a coding including a bit for the sign. 

USER output 

PL - 3.5.6 - 8 

The user end analog output interface shall be according to the Table 3.5-12. 

Measurement range 0 - 5.12 V 

User output impedance < 10 kΩ (1kΩ recommended) 


-14 V 

(tolerance) 


Table 3.5-12 : Analog monitoring output interface (USER side) 

Measurement chain accuracy : 

The End Of Life inaccuracy of the complete measurement chain from the input (including cables and source as 

required in Table 3.5-12) up to and including Analogue/Digital conversion is 0.63 % of the full scale (that is to say 

32 mV with respect to 5.12 V). 

They are listed hereafer in Table 3.5-36. 




PL - 3.5.6 -21 


These active analog interface shall be documented in the instrument units electrical ICD and 

control/command IDS. 




3.5.6.2.2 Thermistor acquisition 

DHU input 

Type of receiver analog input conditioned by DHU 

Pull-up voltage + 10.099 V ( 

Input impedance 6.5 kΩ for Fenwal 

3.01 kΩ for Rosemount 

Input voltage range 0 to 5.12 V 

Resolution 11 bits 

(coded on 12 bits, with 1 bit for sign) 

Absolute accuracy 1.0% not taking into account thermistor inaccuracy 

Differential input capacitance 1 µF 


(tolerance) 



Table 3.5-13 : Electrical characteristics of the DHU TH input interface 

Interconnection 

PL - 3.5.6 - 9 

USER output 

The cable for thermistor acquisition shall be according to Table 3.5-14. 

Cable Twisted shielded 

Return Return signals are grouped 

Return signal connected to secondary ground at DHU 

input 

Table 3.5-14 : Interconnection characteristics 

Type of interface Thermistor 

Thermistor type Fenwal Fw 526-31 BS12-153 (15 kΩ at 25°C) or 

Rosemount 118 MF2000 (2 kΩ at 0°C) 


(tolerance) 



Table 3.5-15 : Output characteristics 

Measurement chain accuracy : 

The End Of Life inaccuracy of the thermistor measurement chain is : 

1.14 % of full scale for Fenwal thermistor 

0.59 % of full scale for Rosemount thermistor 

This inaccuracy do not include thermistor tolerances. 

PL - 3.5.6 -22 

These thermistor acquisition interface shall be documented in the instrument electrical ICD and 

control/command IDS 




3.5.6.2.3 Digital relay acquisition 

DHU input 


Single ended digital receiver, powered by DHU 

Secondary ground used as reference 

Shielding 

Twisted shielded 

Threshold level 

Emission properties (Acq) 

1.5 V ≤ U ≤ 2.5 V 

Voltage 

< 5.5 V 

Current 

0.1 mA to 10 mA 



(tolerance) 

short circuit to secondary ground 

Table 3.5-16 : Electrical characteristics of the DHU DR input interface 

USER output 

PL - 3.5.6 - 10 

Protocol 

The output interface at the user end shall be according to the Table 3.5-17. 

Type of transmitter 

Resistance 

Passive open/closed contacts 

Open 

> 1 MΩ 

Closed 

< 50 Ω 


short circuit to secondary ground 

(tolerance) 


Table 3.5-17 : Digital relay monitoring output interface (USER side) 

Relay status S/W register 

Open 1 

Closed 0 

Table 3.5-17a : DRS protocol in S/W register 

It may be noticed that return lines shall be floating and that each relay shall have a dedicated return (see PL - 4.4.2 - 

7). 

PL - 3.5.6 -23 

These digital relay acquisitions shall be documented in the electrical units ICD and control/command IDS. 




3.5.6.2.4 Digital bi-level acquisition 

DHU input 


Differential (1 return per 2 signals) 

Shielding 

Twisted shielded 

Threshold level 

Input Impedance 

1.47 V ≤ U ≤ 2.51 V 

differential 

“≥ 425 kΩ || ≤ 230 nF (ON, acquisition) 

differential 

≥ 20 MΩ || ≤ 1 µF (ON, outside acq.) 


high input > 400 kΩ (at power ON) 

low input > 25 kΩ 


≥ 1 kΩ || ≤ 1 µF (at power OFF) 



(tolerance) 

-3 V 

USER output 

PL - 3.5.6 - 11 

Table 3.5-18 : Electrical characteristics of the DHU DB input interface 

The output interface at the user end shall be according to the Table 3.5-19 . 

Output voltage 

Low level 

High level 

User output impedance < 10 kΩ 


(tolerance) 

Single ended 

0 V 

3.5 V 

-3 V 


Table 3.5-19 : Digital bilevel monitoring output interface (USER side) 

PL - 3.5.6 -24 

These bilevel acquisitions shall be documented in the units electrical ICD and control/command IDS. 




3.5.6.2.5 16 bits serial acquisition 

PL - 3.5.6 - 12 

The Digital Serial Acquisition (AS8/16) signals (Sample, Data, and Clock) shall comply with the interface 

requirements of the SBDL differential driver/receiver interface, specified hereafter. 

The Standard Balanced Digital Link (SBDL) is a fully differential interface, with a « true line » and a « complementary 

line » (see Figure 3.5-15). The status of the signal is defined as high when the true line has a positive voltage « 1 » 

level with reference to the ground and the complementary line has a « 0 » level with reference to the ground. The 

signal is defined as low when the true line is at « 0 » and the complementary line is at « 1 ». 

DRIVER 

+ 

- 

USER output (data line) 

Figure 3.5-15 : Electrical interface for the SBDL 

DRIVER RECEIVER 

Data USER DHU 

Clock DHU USER 

Enable DHU USER 

Table 3.5-20a: DRIVER and RECEIVER values for Data, Clock and enable 



RECEIVER 











229 Ω < Z < 310 Ω for f ≥ 52 kHz, 

229 Ω < Z < 474 Ω for f ≥ 0 Hz 




(tolerance) 


Table 3.5-20 : Electrical characteristics of the SBDL complementary driver 

+ 

- 

+


DHU input 


.. 


High level 


Low level 



differential 





(tolerance) 


-13 V 


PL - 3.5.6 - 13 

The AS8/16 acquisition timing shall be according to the figures below. The values specified are valid at the 

SBDL output when it is loaded with 1.2 nF ||10 kΩ. 

DS8 Sample 

DS8 Clock 

DS8 Data 

t2 

t1 

t3 

t4 t5 

t7 t8 

t6 

B0 B1 B2 B3 B4 B5 B6 B7 

Figure 3.5-16 : Digital Serial Acquisition (8 bit) timing 



t9


t2 

DS16 Sample 

DS16 Clock 

DS16 Data 

t1 

t4 t5 

t6 

B0 B1 B2 B3 B4 B5 B6 B7 B8 B9 B11 B13 B15 

B10 B12 B14 



t3 

t2+t4 

t7 t8 t9 

t8 t9 

Figure 3.5-17 : Digital Serial Acquisition (16 bit) timing (B0 is MSB) 

AS8 AS16 

t1 64 t ± t as for AS8 


t3 60 t ± t 124 t ± t 


t5 4 t ± 0.1 t as for AS8 

t6 2 t ± 0.25 t as for AS8 

t7 < 16 t as for AS8 

t8 < 1.2 t as for AS8 

t9 > 1.5 t as for AS8 

Rise time 0.2 µs < tr< 1.0 µs as for AS8 

Fall time 0.2 µs < tr< 1.0 µs as for AS8 

where t =2 -20 s ≈ 0.95 µs 

Table 3.5-22 : Digital Serial Acquisition timing 

Note : bit shift shall be done t8 max following falling edge of clock. 

PL - 3.5.6 -25 

These 16 bits serial acquisitions shall be documented in the units electrical ICD and control/command IDS. 

t1


3.5.6.3 Time distribution and synchronization 

3.5.6.3.1 Clock (1 pps) 

PL - 3.5.6 - 14 

The pulse signal (from Platform GPS) characteristics are given here followed 

the clock signal uses RS-422 differential interfaces having the following characteristics : 

Driver output max voltage - 0.5 V to + 6 V 

Driver output signal level loaded ± 2 V 

unloaded ± 4 V 

Driver load impedance 96 to 124 Ohm 

Rise fall time < 400ns when loaded with 1.2nF || 10kOhm 

Receiver input voltage range - 7 V to + 7 V 

Receiver input sensitivity ± 200 mV 

Receiver input resistance 4 kOhm minimum 

In orbit In ground Functional tests AIT 

GPSA FM GPSA EM with GPSA Functional GPSA FM with 

GSSL 

Model (HW PPS) GSSL 

Period 1s 1s 1s 1s 

Period accuracy ± 5µs ± 5µs ± 100µs ± 5µs 

Pulse duration 1µs 1µs 10µs 1µs 

Pulse duration 

accuracy 

PL - 3.5.6 - 15 

±32ns Depends on 

GSSL mode 

The pulse date is transmitted via the MIL STD 1553B bus. The time reference will be on the rising edge of the 

pulse. By convention, the pulse signal is active at high level. 

PL - 3.5.6 -26 

± 5µs 

The pps interface shall be documented in its electrical ICD and control/command IDS. 




3.5.6.4 MIL-STD-1553B bus 


The reference standard is MIL-STD-1553B Notice 2. 

PL - 3.5.6 - 16 

The payload shall use the long stub configuration (transformer-coupled to the bus). 

The BC interface coupler is the DDC BU 61582. 

PL - 3.5.6 - 17 

The RT interface coupler shall be the DDC BU 61582. 

The MIL-STD-1553B bus interface is described in section 3.4.3. 

PL - 3.5.6 -27 

This MIL-STD-1553B interface shall be documented in the units electrical ICD and control/command IDS. 

3.5.6.5 Deleted 




3.5.7 ELECTROMAGNETIC INTERFACE REQUIREMENTS 

3.5.7.1 Conducted Emission & Susceptibility Requirements 

3.5.7.1.1 Conducted emissions on power lines 

The payload shall not generate conducted emissions on the unregulated power bus exceeding the following 

requirements. 

a-Frequency domain (Narrowband) 

PL - 3.5.7 - 1 

The limits given in Figure 3.5-19 apply to payloads which supply or absorb up to 30 W of power. For higher 

absorbed or supplied power levels, the limit is increased by 20×log(P/30) up to the maximum reference 

parameters defined in this figure. 

120 

100 

80 

60 

40 

20 

IdBµAe ff 

Maximum 

P < 30W 

0 

1,0E+0 10,0E+0 100,0E+0 1,0E+3 10,0E+3 100,0E+3 1,0E+6 10,0E+6 100,0E+6 


Figure 3.5-19 : Conducted emission over the power supply bus (Narrowband) 

PL - 3.5.7 - 2 

The measurements shall be carried out in differential mode and common mode. (See test set-up section 

6.1.8.7.1), between 10 Hz and 50 MHz. If the low frequency limit cannot be 10 Hz, but between 10 and 60 

Hz, then a test in time domain is required above 10 Hz (i.e. time domain test cannot be suppressed). 




b- time domain 


PL - 3.5.7 - 3 

A limit of 30 mA peak, read in a bandwidth greater than 75 MHz, is defined for really delivered or absorbed 

power levels less than 30 W. For higher absorbed or supplied power levels, the limit is weighted by a factor 

P/30, yet without exceeding 1 A peak. This limit is applicable to the frequency domain beyond 10 Hz. 

PL - 3.5.7 - 4 

The measurements shall be carried out in differential mode and common mode.(See test set up section 

6.1.8.7.2). 

c-Transient signals (see test set-up section 6.1.8.7.2) 

Turn on transients 

PL - 3.5.7 - 5 

The inrush current shall meet the following requirements: (see Figure 3.5-20) 

i) di/dt < 2.106 A/s 

ii) Imax. * t1 < 400µC and Imax < 20 A 

iii) I < 2*Inom for t1 < t < t2 where Inom = Pmax specified/(Minimum Voltage as defined in PL-3.5.3-6) 

iv) t2 = 50 ms 

C urrents (A m ps) 

Im a x 

2xInom 

In o m 

t 

t1 t2 

Figure 3.5-20 : Inrush Current profile 

Turn off transients at payload switch-off: 

PL - 3.5.7 - 6 

The voltage transients superimposed on the power supply voltage shall be measured in both differential and 

common mode and shall be compliant with Figure 3.5-21. 




PL - 3.5.7 - 7 


The current transients shall remain within the[I Nominal , I = 0 A] range. 

% Power supply line range 

160 

140 

120 

100 

80 

60 

Authorized range 

40 

t1 

t(µs) 

1,00E+00 1,00E+01 1,00E+02 1,00E+03 1,00E+04 

Operational transients 

PL - 3.5.7 - 8 

t2 

t1= 10µs t2= 20 µs t3= 1 ms Vbus = 37 V 

Figure 3.5-21 : Off-switching transient 

The current transient on the power bus shall be less than 2.10 4 A/s. 



t3


3.5.7.1.2 Conducted susceptibility on power lines 

PL - 3.5.7 - 9 

The instrument units shall preserve nominal performance when the following perturbations occur on the 

primary power supply lines. 

a- Sine wave signals 

Differential and common-mode signal injection of a sine signal as defined in Figure 3.5-22. While limits are 

expressed as voltages, injected current shall nevertheless be measured for each test frequency, and shall by no 

means exceed 1 A eff., reducing voltage if needed. 

Above 10 kHz, an amplitude modulation of 30 % at a frequency of 1 kHz shall be superimposed to the 

perturbing signal. 

The test shall comply with methods CS01 and CS02 of MIL-STD-462 with voltage measured between the minus 

wire and the metallic ground for the common mode. 

(See test set up section 6.1.8.7.3). 

Sweep rate for the sine signals shall be less than 1 octave/minute. 

V(Veff) 

1 

0,8 

0,6 

0,4 

0,2 

0 

Differential mode 

Common mode 

1,00E+00 1,00E+01 1,00E+02 1,00E+03 1,00E+04 1,00E+05 1,00E+06 1,00E+07 1,00E+0 

Frequency(Hz) 

Figure 3.5-22 : Susceptibility to sine conducted emissions 

b-Square wave signals 

Square signals : 1 Vpp from 10 Hz to 500 kHz 

The square signal sweep rates shall be less than 1 octave/minute, and the square wave rise time less than 50 ns. 

The measurement shall be carried out in differential mode (See test set up section 6.1.8.7.3). 




c-Transient signal 

The signal (Figure 3.5-23) shall be applied for not less than 1 minute by method CS06 of MIL-STD-462, using a 

positive, then a negative, at a 1 Hz and at a 10 Hz recurrence. 

Such signal shall be applied in differential mode and common mode between the return line and the mechanical 

ground. The level of injected signal is Vbus in differential mode and 12 V in common mode. 

The test shall be performed at Vbus = (Minimum Voltage value as defined in PL-3.5.3-6) and Vbus = 37 V. 

The requirement shall be verified with no need for achieving the specified voltage if the limit current value of 3 A 

peak is measured at the input of the tested unit. 

Source impedance shall be simulated by the LISN as defined in section 6.1.8.5.1 (See test set up section 

6.1.8.7.4). 



percent of nominal voltage 


Rise and fall time in all parts < 200 ns 

Reference level of the amplitude = 100% 

Figure 3.5-23 : Conducted susceptibility, transient wave shape 



Time (µs)


3.5.7.1.3 Susceptibility wrt intermodulation and cross-modulation 

PL - 3.5.7 - 10 

The Payload receiver shall not be perturbed by signal injection as defined hereafter. 

The payload receiver units shall be characterised in terms of response to intermodulation and crossmodulation 

phenomena as well as of rejection w.r.t. spurious signals. Tests shall be run by methods CS03, 

CS04 and CS05 of MIL-STD-462, and they shall be restricted to the frequency bands used by the satellite 

(see section 3.5.7.2.1). 

3.5.7.1.4 Conducted susceptibility of interface signals 

PL - 3.5.7 - 11 

2 

1 

0 

Interface signals shall not be sensitive to common mode voltage as follow : 

• 1.5 V DC, 

• 1 V peak to peak with current limitation at 1 A from 15 kHz to 180 kHz then falling to 0.2 V at 15 MHz. 

The measurement shall be carried out in common mode (see test set up section 6.1.8.7.5). 

V pp 

Common mode voltage 

1,00E+04 1,00E+05 1,00E+06 1,00E+07 1,00E+08 


Figure 3.5-26: Common mode voltage 




3.5.7.2 Radiated Emission and Susceptibility Requirements 

The requirements hereafter exclusively apply to units. In case of excess emission or susceptibility, the contributions 

from harness/wiring and the test facilities have to be determined. 

PL-3.5.7-12 

The measurement shall be carried out up to 1 GHz. For RF equipment, the measurement shall be carried out 

up to 18 GHz. 

3.5.7.2.1 Emissions radiated by E-field 

PL-3.5.7-13 

Emission radiated from payload units shall not exceed the susceptibility level of Platform components units 

and of launch vehicle components. 

a) Platform constraints : 

Platform constraints for emissions radiated by E-field are defined in the following figure. The two particular bands 

with reduced emission concern GPS receiver and TTC receiver which shall not be perturbed by payload emission. 

The electric field, measured at 1 m by method RE02 of MIL-STD-462, radiated both by the test equipment and by 

associated, representative harness/wiring, shall not exceed the limits set in Figure 3.5-24 (mission dependent 

value) 

The measurement range is 10 kHz to 1 GHz in Narrow band except for RF equipment. 

E (dBµV/m) 

90 

80 

70 

60 

50 

40 

30 

20 

10 

0 

1.E+04 1.E+05 1.E+06 1.E+07 1.E+08 1.E+09 1.E+10 1.E+11 


Figure 3.5-24 : Radiated emission, E-field, Narrow band 




Requirements rationale E(dBµV/m) Frequency Band 

GPS 30 1555-1596 MHz(8) 

PROTEUS receiver 30 2025-2110 MHz(10) 

Table 3.5-24 : Requirements about radiated emission, E field and narrow band 

Payload emission will be studied case by case (specific narrow bands due to the payload emission will be 

identified and negotiable). 

b) Launch vehicle constraints : 

As no EIRP (Equivalent Isotropically Radiated Power) is emitted by PROTEUS satellite (payload included) during 

launch phase, there should be no susceptibility problem with the chosen launch vehicle. 

In case of radiated emissions for particular payload, analysis shall be conducted to appreciate the compatibility 

with launch vehicle receivers. 




3.5.7.2.2 Radiated electric susceptibility 

PL-3.5.7-14 

The Payload shall be free from any misfunctionning or performance degradation when subjected, by method 

RS03 of MIL-STD-462, to electrically generated electromagnetic radiation within the limits specified in terms 

of electric fields on Figure 3.5-25. 

In the case of receiver units, this test is not applicable within their receiving band. 

To that effect, the above field shall be 50 % amplitude-modulated by a sinusoid of a frequency like those frequencies 

at which the unit was found conduction-susceptible. The carrier-to-modulating frequency ratio shall be more than 5. 

Susceptibility shall be tested up to 1 GHz. Susceptibility testing of the RF units shall run up to 18 GHz, as applicable 

up to a higher frequency for specific units, with functional emission frequencies used as the test frequencies for all 

units. 

Sine wave sweep rate shall not exceed 1 octave/minute, and the sine signal shall be amplitude-modulated by a 

square signal with a 30% modulation rate in the dedicated (e.g. radar) bands. 

20 

15 

10 

5 

Radiation Origin E(V/m) Frequency Band Applicable for 

PROTEUS emitter 14 2200-2290 MHz any equipment 

E(V /m ) 

2200MHz-2290MHz 

14V/m 

0 

1,00E+04 1,00E+05 1,00E+06 1,00E+07 1,00E+08 1,00E+09 1,00E+10 1,00E+11 


Figure 3.5-25 : Radiated susceptibility, E-field 

The following tables show the levels generated by launch vehicles telecommands on satellite equipment units, these 

levels are given at the satellite / launch vehicle interface. 




PL-3.5.7-15 


All the equipment shall withstand this environment without permanent performance degradation. 

DELTA II : 

DELTA II radiated field levels: 

Frequency (MHz) 2241.5 2244.5 5765 

E (dBµV/m) at 1m 140 140 152 

RF environment on Vandenberg Air Force Base launch site: 

ROCKOT : 

14 KHz to 5762 MHz E = 140 dBµV/m 

5762 MHz to 5768 MHz E = 152 dBµV/m 

5768 MHz to 40 GHz E = 140 dBµV/m 

Frequency (MHz) Max Antenna E-field (dBµV/m) 

power (dBW) With fairing Without fairing 

120-130 12.3 107 119 

1040-1050 8.0 105 117 

1015-1025 6.8 100 112 

2700 – 3000 (tracking) 20.0 (pulsed 

mode) 

107 119 

Table 3.5-28 ROCKOT L/V transmitters radiated field levels 

Figure 3.5-31: ROCKOT Launch Vehicle RF environment 




SOYUZ : 


Frequency Band (MHz) LV Field Intensity (dBµV/m ) 

0.014 - 200 60 

200 - 250 150 

250 - 380 60 

380 – 620 80 

620 - 680 140 

680 - 970 80 

970 - 1060 140 

1060 - 1250 80 

1250 - 2700 100 

2700 - 3000 145 

3000 - 3300 100 

3300 - 3500 145 

3500 - 10000 100 

>10000 85 

Table 3.5-29: LV and launch base mission Spectra (Soyuz ST configuration) 




Figure 3.5-32: LV and Launch base Emission Spectra (Soyuz ST Configuration) 

More precise and detailed information is available in Launch vehicles User Manual. In project phase B, these 

characteristics are checked to make the satellite compliant with the launch vehicle. 




3.5.8 ESD PROTECTION 

3.5.8.1 Direct arc discharge 

PL - 3.5.8 - 1 

No malfunction shall occur when the payload is submitted to a direct repetitive arc discharge of at least 10 

mJ energy. 

Figure 3.5-27: Unit under direct arc discharge 

3.5.8.2 Indirect arc discharge 

PL - 3.5.8 - 2 

No malfunction shall occur when the payload is submitted to an indirect repetitive arc discharge of at least 

10 mJ energy. 

300 mm 

Figure 3.5-28: Unit under indirect arc discharge 




3.5.9 MAGNETIC FIELD INTERFACE REQUIREMENTS 

3.5.9.1 Emission requirements 

PL - 3.5.9 - 1 

The total payload magnetic moment shall be lower : 

• 1 A.m² .when the Payload is OFF 

• 5 A.m² when the Payload is ON. 

PL - 3.5.9 - 3 

The Payload Supplier shall first identify all ferromagnetic material and the amounts used in the fabrication of 

its flight hardware. 

Secondly, the Payload Supplier shall provide data on residual magnetic dipoles of its flight hardware to the 

Satellite Contractor for incorporation into the overall magnetic budget. 

3.5.9.2 Susceptibility requirements 

PL - 3.5.9 - 2 

The payload shall withstand a magnetic field created by the PROTEUS platform of up to: 

• 1 Gauss in satellite nominal mode excepted in the volume C1 (co-ordinates indicated in Table 3.5-25) 

where its value is between 1 Gauss and 3 Gauss 

• 5 Gauss in satellite Safe Hold Mode excepted in the volume C2 (co-ordinates indicated in Table 3.5-26) 

where it can reach 23 Gauss punctually. 

X * (meter) 0 0 0 0 +0.1 0.1 +0.1 +0.1 

Y *(meter) -0.45 +0.45 -0.45 +0.45 -0.45 +0.45 -0.45 +0.45 

Z *(meter) -0.15 -0.15 -0.6 -0.6 -0.15 -0.15 -0.6 -0.6 

* X, Y, Z are payload axes, X (payload) = X(satellite) - 1.07 (in meter) 

Table 3.5-25 Volume C1 where the magnetic field is between 1 and 3 Gauss in satellite nominal 

mode 

X *(meter) 0 0 0 0 +0.2 +0.2 +0.2 +0.2 

Y *(meter) -0.6 -0.6 0.6 0.6 -0.6 -0.6 +0.6 +0.6 

Z *(meter) -0.15 -0.6 -0.15 -0.6 -0.15 -0.6 -0.15 -0.6 

* X, Y, Z are payload axes, X (payload) = X(satellite) - 1.07 (in meter) 

Table 3.5-26 Volume C2 where the magnetic field can reach 23 Gauss in satellite SHM 

These values are the maximum values of the generated magnetic field at PL level. These fields are generated by MTB. 

Note 1 : During normal mode, MTB are used for reaction wheel momentum management. The magnetic field 

induced by MTB use is mission dependent. The value indicated in this specification correspond to 20 Am² MTB use. If 

the MTB use is greater than this 20 Am², a specific analysis shall be carried on at the beginning of satellite phase C 

in order to evaluate the new magnetic field. 

Note 2 : During satellite Safe Hold Mode where MTB are rated at 180 A.m², this magnetic field may reach 5 Gauss 

(except in the volume C2 where it can reach 23 Gauss) but the payload will be OFF. It can be noticed when the 

satellite is switched to Safe Hold Mode, all the payload units are switched OFF (except the 2 lines «8» and «16» that 

may be ON according to their state before SHM) 




The maximum variation (PL - 3.5.9 - 4) will be evaluated probably at the beginning of satellite phase C. This 

maximum variation also occurs during SHM. Nonetheless, as stated before, during normal mode, some variation will 

also occur. 

Moreover, it may be noticed that the generated magnetic field is not uniform inside the whole payload and that the 

maximum value is obtained near the –Ys and -Zs payload faces (at the PF/PL interface of course). 

PL - 3.5.9 - 4 

The payload shall withstand a magnetic field variation created by the PROTEUS platform of up to : 

• 19 Gauss/s in satellite nominal mode excepted in the volume C1 (co-ordinates indicated in the table 

3.5-25) where the maximum magnetic field variation can reach 50 Gauss/s. 

• 94 Gauss/s in satellite Safe Hold Mode excepted in the volume C2 (co-ordinates in the table 3.5-26) 

where the maximum magnetic field variation can reach 430 Gauss/s punctually. 




3.6 STAR TRACKER ASSEMBLY ACCOMMODATION 

3.6.1 GENERAL 

Two Star Trackers are accommodated as a single Assembly. This STA is part of the Attitude and Orbit Control System 

(AOCS) functional chain and will be implemented on the Payload during the Satellite AIT, after the payload mating. 

This section describes the standard STA. Some changes may occur for each mission (orientation angle, mass, 

interface,...) but the main characteristics will be unchanged. 

Consequently, this section provides to the Payload Supplier all the necessary preliminary data and requirements to 

implement the STA on the Payload structure. 

PL - 3.6.1 - 1 

The STA shall be accommodated on the Payload by the Payload Supplier but this accommodation has to be 

agreed by the Satellite Contractor and shall be in accordance with the specifications defined in section 3.6.2. 

The STA will be integrated on the Payload during Satellite AIT, in Satellite Contractor facilities. 

The angle between the mounting plane and the line of sight shall be agreed by the Payload Supplier and the Satellite 

Contractor. 

The Platform Contractor is in charge of the STA thermal control. 

Figure 3.6-1 gives a global view of the standard Star Trackers Assembly. 

Figure 3.6-1 : Standard Star Trackers Assembly 

The standard STA Interface Data Sheet (IDS) is provided in appendix C and STA main characteristics are also given 

in the following sections. 

This standard STA is compatible with a PL interface in carbon. Indeed, this design is limited in case of PL interface in 

aluminium or any other material with a high thermal expansion coefficient to a restricted thermal gradient 

This compatibility shall be assessed on each mission as requested in PL-3.6.1-1. 




This STA may be adapted on amission by mission basis: STR elevation & azimuth, STA radiator panel,...For 

example, concerning adaptation to an aluminium PL interfae, an adapted design was developed to cope with high 

thermal gradient between PL and STA. 




3.6.2 MECHANICAL SPECIFICATIONS 

3.6.2.1 Interfaces 

PL - 3.6.2 - 1 

The Payload shall comply with the interface plane and points described in Figure 3.6-2. 

Figure 3.6-2 : Standard Star Trackers Assembly interface plane 

The STA reference frame is shown in Figure 3.6-2 and defined in section 1.4. Figure 3.6-2b shows interface cross 

section. 




Note A is th e plane defined by th e 8 co n ta ct a rea s 

. The global flatness is 

. Each contact is // to the other according 



0.1 

Figure 3.6-2b : Interface cross section 

// 0.4/40


3.6.2.2 Physical characteristics 


3.6.2.2.1 Mass 

The maximum mass of the STA is 12.50 kg. 

3.6.2.2.2 Volume 

The volume of the standard STA is given in Figure 3.6-3. 

3.6.2.2.3 Centre of gravity 

STR1 

STR2 

Figure 3.6-3 : Standard Star Trackers Assembly volume 

Optical cube STR1 

Optical cube STR2 

The centre of gravity of the standard STA (expressed with respect to its interface see Figure 3.6-2) is provided in 

appendix C. 

3.6.2.2.4 Moments of Inertia 

The moments of inertia of the standard STA, expressed with respect to the STA centre of gravity, are provided in 

appendix C. 

3.6.2.2.5 Stiffness 

Note: Considering that the STA definition is mission dependant, this stiffness above can vary specially if the STA 

definition change is related to the baseplate material. 

The STA first mode frequency is higher than 141,8 Hz. Moreover, main modes are : 

Along X STA axis : 141,8 Hz 

Along Y STA axis : 181,6 Hz 

Along Z STA axis : 347,8 Hz 




A STA FEM simplified model will be provided to the Payload Supplier in a Nastran version 70 format. 

The model will include at least: 

mass, 

location of the CoG, 

moments of inertia, 

modal characteristics. 




3.6.2.2.6 Material 

PL - 3.6.2 - 2 

The material used at STA interface shall be compatible with Carbon (STA honeycomb structure face sheets), 

Permaglass (STA thermal insulating washers) and titanium alloys (STA screws). This applies to all PL material 

located under the STA baseplate area. 

The use of mercury, cadmium and Zinc is prohibited. 

3.6.2.2.7 Alignment 

The angle between each STR line of sight and the "Payload reference line of sight" (line of sight of one of its 

instruments) is mission dependant. 

PL - 3.6.2 - 3 

After payload mechanical assembly mounting, the targeted orientation of the STA reference frame with 

respect to the payload reference frame (Fp) shall be achieved with a maximum deviation of 0.25° (3σ) about 

each axis in Fp. 

Nota: Verification of alignment accessibility to STR optical cube will be done after reception of PL mature CAD model 

during satellite phase B. 

The positions of the STR 1 and STR 2 optical cubes (i.e 5 polished facets) are defined in Figures 3.6-3, 3.6-3c and 

3.6-3d. 

Figure 3.6-3c : F1 view of the STR 1 with reference cube orientation 




Figure 3.6-3d : F2 view of the STR 2 with reference cube orientation 

PL - 3.6.2 - 4 

The 3 axes alignment of each STR with respect to the satellite reference optical cube shall be directly 

measurable at any time during the satellite integration. That is to say that STR optical cubes shall be 

accessible from 2 perpendicular directions in the horizontal plane (no payload appendices, no interference). 

3.6.2.2.8 Geometrical constraints 

PL - 3.6.2 - 5 

The STA shall be accommodated on the « mission dependent » side of the Payload. 

PL - 3.6.2 - 6 

The standard theoretical elevation angle between the STR boresight (Z STR1 and Z STR2) and the STA interface 

plane is 45° ± 2.5°. 

If no accommodation is possible with this standard value, specific studies will be performed and the bracket will be 

modified. 




PL - 3.6.2 - 7 


The accommodation of the STA shall be such that : 

• the theoretical azimuth angle (angle measured in the interface plane between Xp and the projected STR 

boresight (YSTA if there are no washers)) is equal to «mission dependent value». 

The real azimuth angle shall remain compatible of PL-3.6.2.3. 

Figure 3.6-4 : Azimuth and elevation definition : CALIPSO (111° and 45°) 




PL - 3.6.2 - 11 


Moreover, the stability of angles between STA reference frame and payload boresight shall be consistent with 

the satellite pointing knowledge requirement knowing that it impacts directly this budget. 

This stability shall include all in-orbit events such as : 

• Launch bias (launch loads effects, internal mechanical shift (1g to 0g), hygroelastic effects….) 

• Thermo effects… 

PL - 3.6.2 - 8 a 

The STR is functional only if a clearance angle is respected between its line of sight and an incident light 

beam from the Earth or the Sun and also between its line of sight and any satellite (including payload) 

appendages. 

The clearance angles shall be equal or higher than : 

• 40 deg between the Sun and each STR line of sight 

• 32 deg between the Earth and each STR line of sight 

• 40 deg between the satellite appendices and each STR line of sight. 

Note : For inertial mission (as Corot), the possible Moon effects on long duration will be studied. 

3.6.2.3 Dynamic Environment 

3.6.2.3.1 Quasi static acceleration loads 

PL - 3.6.2 - 9 a 

The STA shall not see Quasi-Static loads greater than those defined in Table 3.6-1. 

Axis qualification level (g) 

STA Z axis ± 20 

⊥ to STA Z axis ± 20 

Table 3.6-1 : Quasi static acceleration loads 

Based on delivered STA finite element model (which includes the STR restitution points) and based on these values at 

STR level, the Payload Supplier will be able to adequately define the structure supporting the STA by avoiding 

coupling between these structures. If necessary, the Payload Supplier will define the required notching level during 

payload sine vibration tests. As required in section 4.2.5.3, this notching request will be analysed by the satellite 

contractor and submitted to its approval. 

These levels are also given for the calculation of the maximum forces and the moments at the STA interface. 

3.6.2.3.2 Random vibrations 

PL - 3.6.2 - 12 a 

The random vibrations qualification levels seen at each STA interface along each axis shall be less than the 

spectrum given in Table 3.6-2 and Figure 3.6-6. 




Excitation axis Frequency range Power spectral 

(Hz) 

density (PSD) (TBC) 

Longitudinal(XSTA) 20 0.002 g2 /Hz 

70-120 0.07 g2 /Hz 

140-180 0.02 g2 /Hz 

190-250 0.05 g2 /Hz 

350-500 0.002 g2 /Hz 

600-800 0.03 g2 /Hz 

2000 0.0002 g2 /Hz 

Lateral (YSTA & ZSTA) 20 0.002 g2 /Hz 

70-120 0.07 g2 /Hz 

130-220 0.025 g2 /Hz 

230-300 0.015 g2 /Hz 

350-600 0.04 g2 /Hz 

1000 0.001 g2 /Hz 

2000 0.0002 g2 /Hz 

Table 3.6-2 : Maximum random vibration levels at Star Trackers Assembly level 



Random level [g²/Hz] 


1 

0.1 

0.01 

0.001 

0.0001 

10 100 1000 10000 

Random level [g²/Hz] 

1 

0.1 

0.01 

0.001 


STAyz max level [g²/Hz] 

0.0001 

10 100 1000 10000 


STAx max level [g²/Hz] 



STAyz max level [g²/Hz] 

Figure 3.6-6: Maximum random vibration levels at Star Trackers Assembly level 

For information, and in case of non compliance with the previous requirement at STA interface level, the analysis of 

a deviation request on PL-3.6.2-12 will be based on the analysis of the predicted levels at STR level itself.


3.6.2.3.3 Shock 

PL - 3.6.2 - 10 

The STA shall not see a Shock Response Spectrum greater than the one given in Table 3.6-3. 

Frequency 

qualification level 

(Hz) 

(g) 

100 5 

2000 600 

10000 600 

Table 3.6-3 : Maximum Shock levels at Star Trackers Assembly level 




3.6.3 HARNESS CONSTRAINTS 

PL - 3.6.3 - 1 

The power and signal harness will be provided by the Platform Contractor and delivered to the Payload 

Supplier for accommodation during Payload AIT. 

The thermal control harness shall be provided by the Payload Supplier. 

Figure 3.6-5 summarises wiring interfaces between STA (STRs & H20 connector bracket) and H01, H02 & H03 

connectors brackets. 

Separation 

of the wires 

J08 

Nominal 

STR 

Acq & Cmd 

H02 

STA 

J10 

Thermistors 

for thermal 

control 1 

"Acquisition & Command" 

P/L I/F bracket 

STR 1 STR 2 

H20 

P01 

Nominal 

Thermal 

Control 

Heater 

2 heaters & 3 thermistors 

P05 

P04 

STR1 

power 

P05 

STR2 

power 

H01 

"Power" bracket 

P08 

Redundant 

Thermal 

Control 

Heater 

Towards 

payload (for 

active thermal 

control) 

J08 

Redundant 

STR Acq & 

Cmd 



J05 

H03 

J10 

Thermistors 

for thermal 

control 2 

"Acquisition & Command" 

P/L I/F bracket 

Figure 3.6-5 : STAs wiring (STRs and CTA) 

As described in PL-3.6.3-1, the power (red) and signal (blue) harness will be provided by the platform contractor. On 

the contrary, the thermal control harness (green and orange) shall be provided by the Payload Supplier because this 

harness is common with the rest of the payload thermal control (see the separation of the wires on the Figure: line 

dedicated to the STA active thermal control). 

All the information related to Figure 3.6-5 are given either in section 3.6.3 either in appendix B (PL connectors and 

pin description). 

With respect to all the previously described constraints, the baseline is the following one.


PL - 3.6.3 - 2 a 


With regard to thermal control harness, the payload supplier shall define the wiring routing, shall 

accommodate this harness and shall provide the STA wires with connector P05 up to the H20 bracket fixed 

on the STA and with the good length (length between H20 bracket and Anchor # 2 = 242 mm, length 

between STA Anchor # 2 and H02/H03 = defined by the Payload Supplier). 

The Platform J05 connector on the H20 bracket is DEMA-15S (HD) type. 

The localization (on the STA Frame) of Anchor# 2 is shown on the Figure 3.6-7. 

The satellite contractor will then perform heaters and thermistors connection to STR's during satellite AIT. These 

heaters and thermistors are on the STA and are provided by ALCATEL. 

PL - 3.6.3 - 3 a 

With regard to power and signal, the payload supplier shall define the harness routing from H01, H02/H03 

brackets up to Anchor#1 and Anchor#2 (mounted on the STA baseplate) and accommodate this harness 

according to satellite following requirements : 

• power and signal harness shall be separated by at least 100 mm, 

• bend radius shall be higher than 33 mm (around 3 times the outside diameter), 

• harness routing definition shall be approved by the satellite contractor 

The localization (on the STA Frame) of Anchor# 1 and Anchor# 2 is shown on the Figure 3.6-7. 

On this Figure, J1 = Acquisition connectors and J4 = Power connectors. 




Anchor # 1 Anchor # 2 

STR 1 L to J1 = 250 mm 

STR 1 L to J4 = 275 mm 

H20 Bracket (CTA) 

L to J05 = 242 mm 

STR 2 L to J1 = 605 mm 

STR 2 L to J4 = 575 mm 

STA CTA 

L = 242 

Groung Stud 

Figure 3.6-7 : Anchor# 1 and Anchor# 2 positions on the STA baseplate 




Characteristics of these power and signal wires are given hereafter: 

• Length 

Cables length from STR to H01, H02/H03 brackets are given in the Table 3.6-4. 

CABLES STR1 P01 

H02 J08 

Length from STR to 

anchor* (mm) 

Length from anchor to 

bracket (mm) 

STR1 P04 

H01 P04 

STR2 P01 

H03 J08 



STR2 P04 

H01 P05 

250 275 605 575 

TBD (mission 

dependant) 

Total (mm) TBD (mission 

dependant) 

TBD (mission 

dependant) 

TBD (mission 

dependant) 

* Anchor#1 and anchor#2 are mounted on the STA baseplate 

Table 3.6-4: STR cables length 

TBD (mission 

dependant) 

TBD (mission 

dependant) 

TBD (mission 

dependant) 

TBD (mission 

dependant) 

• Diameter of signal cable = 11 mm 

• Diameter of power cable = 9 mm 

• Overall maximum mass : 

The owerall maximum mass of the 4 STR wires (2 cables signal (N+R) and 2 cables power (N+R) including 

connectors, locking, shielding) is 1.2 kg. 

Note : Considering the STA definition, the STA position on the payload and the STR cables routing on the payload 

are mission dependant, this owerall maximum mass is an allocation mass to take into account on the equipped 

payload mass calculation (see PL-3.1.1-1 specification). The maximum real mass shall be estimated when the real 

total length of these 4 cables will be defined (in order to keep the maximum real mass under this allocated mass, the 

STR cables routing on the payload will goals not to exceed 2.5 m for the 2 cables signal and 3 m for cables power). 

Position of STR connectors is shown in appendix C and these connectors are included in the delivered CAD model. 

3.6.4 THERMAL DESIGN AND INTERFACE REQUIREMENTS 

The two Star trackers are located inside one enclosure (STA) thermally actively controlled at satellite level by a 

redounded heating line driven by the DHU. All STA thermal control items (MLI, radiators, heaters, thermistors, 

insulating washers) will be determined and supplied by the Platform Contractor 

The global thermal conductive coupling between the STA and the payload will be lower than 0.04 W/°C and 

obtained thanks to the insulating washers shown Figure 3.6-2b. 

Moreover, the STA will be radiatively decoupled from the payload (MLI covering all sides except the radiative areas). 

The equivalent efficiency of the ZSTA MLI blanket (MLI between STA and payload) is 0.1 W.m2 /°C. 

The heat rejection capability of the STA is achieved by a radiative area (constituted of silvered SSM) located on the 

opposite side of the mounting plane to avoid any radiative coupling with the payload. 

PL - 3.6.4 - 1 a 

The tension generated by each of the 8 interface screws (provided by the payload Supplier) shall be higher 

than 7400 N and less than 12600 N.


PL - 3.6.4 -2 


The screws shall be in ISO Standard (M5), the tapping material shall be Aluminium (insert type) and the 

mechanical strength of both screws and tapping permit to apply these tensions. 

PL - 3.6.4 -3 

The Payload supplier shall define the tightening torque of the screws in order to respect the values specified 

and shall take into account,when this tightening torque is applied, the specific requirements in term of 

tightening procedure in the presence of thermal washers, as indicated in the Appendix D. 

The Payload supplier shall provide justification associated (validated tightening procedure for example). 

PL - 3.6.4 -4 

Each Payload insert supporting the STA shall be capable to withstand mechanical loads : 

• Normal load : 2450 N 

• Shearing load : 7830 N 

Nota: In order to realize the analysis the payload supplier shall provide: 

the temperature range [min,max] of the payload interface on each Payload thermal case 

the characteristics of the payload interface (material, expansion coefficient, stiffness) 

a payload panel thermo-elastic model (supporting the STA) would be appreciated. 

3.6.5 CLEANLINESS REQUIREMENTS 

N.A 

3.6.6 STA GROUNDING ON PAYLOAD 

PL - 3.6.6 -1 

The STA shall be grounding on payload via the two ground braids provided by the Platform contractor (i.e. 

ground braids are STA parts). 

The PL supplier shall connect these 2 ground braids with 2 ground studs (for STA Grounding redundancy) 

present on payload side. 

If these two ground braids can't be connected with the Payload Grounding Point (i.e. 2 ground studs as 

indicated on § 4.2.2.2), the payload supplier shall foresee 2 dedicated ground studs as shown on Figure 

3.6-8 (TBC values are typical values which shall be defined by the Payload Supplier depending on payload 

design). 

The ability to use Payload Grounding Point for STA grounding can be discussed with the Satellite Contractor. 

If the 2 dedicated ground studs are needed, they shall be located near the STA ground stud at a distance 

about 100 mm. 

The localization (on the STA Frame) of the STA ground stud is shown on the Figure 3.6-7. 

The localization of these 2 dedicated ground studs shall be identified in the payload ICD. 




(TBC) 

(TBC) 

(TBC) 

(TBC) 

Figure 3.6-8: Ground stud configuration for STA grounding 




3.7 GROUND SUPPORT EQUIPMENT INTERFACES 

General requirements on environmental conditions, design rules, verification and tests are provided hereafter for all 

GSE that will be used in ALCATEL facilities and/or will be «in contact» with the platform. 

All GSE including OGSE, TGSE… shall be provided with similar data described hereafter. 

3.7.1 MECHANICAL GSE INTERFACES 


PL - 3.7.1 - 1 

The payload shall be provided with adequate MGSE mechanical interface points for ground handling, lifting 

and transportation, in all intermediate assembly configurations before the Payload/Platform fitting. The 

payload shall also provide the associated MGSE. In a more general way, any item with a mass > 10 kg shall 

be provided with adequate MGSE mechanical interface point for handling, lifting and integration. The 

payload shall also provide the associated MGSE. 

The I/F point between MGSE and unit shall be designed fail-safe. 

The Satellite handling will be directly performed through platform dedicated handling attached fittings. (see 

paragraph 4.2.2.4). 

3.7.1.2 Requirements for delivered MGSE 

The MGSE delivered to ALCATEL for payload specific operations before PL/PF fitting shall comply with the following 

specific requirements. 

3.7.1.2.1 Environment 

3.7.1.2.1.1 Operational climatic environment 

PL - 3.7.1 - 2 

The MGSE shall operate in the following ranges : 

• temperature between +10° C and +40° C 

• hygrometry between 20% and 80% 

• pressure equivalent to the sea level 

Usually, the MGSE will be used in clean rooms with the following environment : 

particular cleanliness class 100000 or 10000 in case of presence of optics 

molecular cleanliness 10 -6 g/cm 2 

temperature 22° C ±3° C 

hygrometry 30 < HR (%) < 60 

pressure ambient 

3.7.1.2.1.2 EMC 

PL - 3.7.1 - 3 

The MGSE shall be protected against possible EMC disturbances. 




3.7.1.2.1.3 Non operational mechanical environment 

PL - 3.7.1 - 4 

The MGSE shall be mounted and packed so as to withstand shocks and vibrations of handling and 

transportation as defined herein: 

• Vibration (road and air) : 5.5 - 200 Hz +/- 1.5 g 

• Shocks (road and air) : up to 8 g for 5-50 ms 

• Acceleration (air) : up to 3 g constant vertical (banking). 

3.7.1.2.1.4 Non operational thermal environment 

PL - 3.7.1 - 5 

During transportation, the GSE container shall withstand the following environment 

• temperature -40° C < T < +50° C 

• hygrometry 1 < HR (%) < 100 

3.7.1.2.1.5 Non operational pressure environment 

PL - 3.7.1 - 6 

During transportation, the GSE container shall withstand the following environment 

• pressure 200 hPa 

• pressure drop speed 143 N/m².s 

3.7.1.2.2 Design requirements 

3.7.1.2.2.1 General design rules 

PL - 3.7.1 - 7 

The critical requirements are 

• modular design using standard elements 

• minimisation of the hazards towards personnel 

3.7.1.2.2.2 Electrical design rules 

PL - 3.7.1 - 8 

The electrical elements of the MGSE shall be designed for a main power supply of 220 V ±10% single phase, 

50 Hz ±20%. 

PL - 3.7.1 - 9 

The MGSE shall be grounded via a single point attachment. 




3.7.1.2.2.3 Mechanical design rules 

PL - 3.7.1 - 10 

MGSE shall be designed applying safety factors of Table 3.7-1 and environment given in section 5.11.2. 

Yield Ultimate 

MGSE 2 3 

Lifting device 3 5 

3.7.1.2.3 Verification and tests requirements 

3.7.1.2.3.1 Design file 

Table 3.7-1 : Factors of safety 

The design file mainly consists in drawing associated to each element of the MGSE. 

PL - 3.7.1 - 11 

This design file shall be approved before the beginning of the manufacturing 

3.7.1.2.3.2 Justification file 

Each element shall be justified by analysis. All these analyses shall be recorded in the justification file. 

PL - 3.7.1 - 12 

This justification file shall be approved before the beginning of the manufacturing 

3.7.1.2.3.3 Verification and test matrix 

PL - 3.7.1 - 13 a 

The delivered MGSE shall be submitted to at least : 

• Weighing 

• Visual inspection 

• Initial and periodic static test, the period is fixed according to the MGSE criticality (the proof loads shall 

be the double of the maximum required load) 

• Functional tests 

• Waterproofness test 

• EMC tests 

• Annual inspection report from a nationally recognized testing organism; validity period shall cover 

launch campaign including six months launch delay. 




3.7.1.2.4 Required documentation 

PL - 3.7.1 - 14 

The documentation to be delivered with the hardware give design description and functional explanation of 

each equipment. The supplier shall deliver an Acceptance Data Package (ADP) containing at least the 

following documents: 

• Technical manual 

• User manual 

• Acceptance test procedure and programs. A supplier acceptance test report shall be provided including 

the test and operations which have been carried out on equipment or components 

PL - 3.7.1 - 15 a 

Upon level acceptance, the following additional documentation shall be delivered: 

• log book, 

• acceptance report, 

• a certificate of conformity «CE» if the MGSE come from Europe, 

• a certificate of conformity according to US regulations if the MGSE come from US. 

Any missing part of the ADP shall cause the acceptance process to be stopped. 

The ADP documents shall be delivered in two copies, plus one for each set of delivered equipment. 




3.7.2 ELECTRICAL GSE INTERFACES 


3.7.2.2 Requirements for delivered EGSE 

The EGSE delivered to ALCATEL for payload specific operations at satellite level shall comply with the following 

requirements. 

3.7.2.2.1 Environment 

3.7.2.2.1.1 Operational climatic environment 

PL - 3.7.2 - 1 

The EGSE shall operate in full compliance with their requirements in the following environment : 

• temperature between +10° C and +40° C 

• hygrometry between 30% and 80% 

• pressure ground level to 2000 m 

Usually, the EGSE will be used in clean rooms with the following environment : 

particular cleanliness class 100000 or 10000 in case of presence of optics 

molecular cleanliness 10 -6 g/cm 2 

temperature 22° C ±3° C 

hygrometry 30 < HR (%) < 60 

pressure ambient 

3.7.2.2.1.2 EMC 

PL - 3.7.2 - 2 

The EGSE and cables to be used during functional testing shall be designed under the following guidelines : 

• low susceptibility to external interference (conducted interference trough signal and power lines, radiated 

interference) 

• low conducted emission and radiated emission to avoid interference with satellite and other test 

equipment 

• the use of the equipment shall not introduce ground loops 

A 6 dB margin for the EGSE shall be taken below the maximum values defined in Figure 3.5-19 and Figure 

3.5-24. 




3.7.2.2.1.3 Non operational mechanical environment 

PL - 3.7.2 - 3 

The EGSE shall be mounted and packed so as to withstand shocks and vibrations of handling and 

transportation as defined herein: 

• Vibration (road and air) : 5.5 - 200 Hz +/- 1.5 g 

• Shocks (road and air) : up to 8 g for 5-50 ms 

• Acceleration (air) : up to 3 g constant vertical (banking). 

3.7.2.2.1.4 Non operational thermal environment 

PL - 3.7.2 - 4 

During transportation, the EGSE container shall withstand the following environment 

• temperature -40° C < T < +50° C 

• hygrometry 1 < HR (%) < 100 

3.7.2.2.1.5 Non operational pressure environment 

PL - 3.7.2 - 5 

During transportation, the EGSE container shall encounter the following environment 

• pressure 200 hPa 

• pressure drop speed 143 N/m².s 

3.7.2.2.2 Design requirements 

3.7.2.2.2.1 Protection 

PL - 3.7.2 - 6 

This covers resistance against moisture, salts, corrosion, fungus. 

Hygroscopic materials (e.g. wood) and components shall not be used for preservation, casting or similar 

protection. 

Corrosion sensitive materials are to be avoided or shall be provided with an appropriate high quality surface 

tempering and/or finish for the specified environmental conditions. 

3.7.2.2.2.2 Electrical design requirements 

PL - 3.7.2 - 7 

EGSE equipment shall be designed for a main power supply of 220 V ±10% single phase, 50 Hz ± 20%. 

PL - 3.7.2 - 8 

The maximum current demand from the AC power lines shall be less than 120% of the maximum static input 

current. 




PL - 3.7.2 - 9 


The tolerable noise and ripple level on the AC power line shall be 5%. 

PL - 3.7.2 - 10 

The connectors types shall be 220 V NF USE. 

PL - 3.7.2 - 11 

Fuses or circuit breakers shall be implemented on all main inputs of the power lines. 

PL - 3.7.2 - 12 

The insulation between any output terminal and the AC power line shall be higher than 10 Mohm. 

PL - 3.7.2 - 13 

The EGSE grounding concept shall be the following : 

• All signal lines are floating ones with respect to spacecraft I/F 

• To respect galvanic insulation, the electronic components of the EGSE side must be referenced to EGSE 

ground 

• The umbilical signals which are referenced to the payload primary ground (respectively to the secondary 

ground) shall be processed in EGSE by electronic unit referenced to the payload primary ground 

(respectively to the payload secondary ground) and shall be isolated from EGSE ground, [analysis] 

• The EGSE shall be grounded via a single point attachment. 




3.7.2.2.3 Verification and tests requirements 

PL - 3.7.2 - 14 

The acceptance tests will be done in nominal operating environment. 

PL - 3.7.2 - 15 

The necessary calibration and maintenance of the above equipment shall be provided. 

The subcontractor of the testing EGSE equipment shall provide a document in which can be found the 

suggested solution for each requirement. 

PL - 3.7.2 - 16 

The acceptance tests shall be performed according to an approved acceptance test procedure. 

This shall include at least: 

• A visual inspection of hardware 

• Verification that EGSE equipment design meets the requirements 

• Identification of defects in material or workmanship 

• Identification of unexpected interference between assemblies 

• Compatibility verification to interfacing equipment, in particular with the platform or spacecraft 

equipment 

• Incorporate a review of the log book and required documents 

• An acceptance test for each model built 

• Qualification of GSE to be used in working environments. 

• Verification of hazardous order inhibition shall be performed 

There will be an acceptance test for each model built. 

The EGSE system design report will include verification matrixes to define EGSE acceptance tests according to 

the design and requirements. 




3.7.2.2.4 Required documentation 

PL - 3.7.2 - 17 

The documentation to be delivered with the hardware give design description and functional explanation of 

each equipment. 

The supplier shall deliver an Acceptance Data Package (ADP) containing the following documents : 

• Equipment technical specification (update) 

• Manufacturers handbook of all commercial equipment 

• Technical manual 

• User manual 

• Acceptance test procedure and programs. A supplier acceptance test report shall be provided including 

the test and operations which have been carried out on equipment or components 

PL - 3.7.2 - 18 a 

Upon level acceptance, the following additional documentation shall be delivered: 

• log book, 

• acceptance report, 

• a certificate of conformity justifying that: 

- protective devices are available on EGSE primary circuits 

- no live part is accessible to personnel, 

• a certificate of calibration with the date of validity. 

Any missing part of the ADP shall cause the acceptance process to be stopped. 

The ADP documents shall be delivered in two copies, plus one for each set of delivered equipment. 





Chapter 5 : Payload environment requirements 




Status 

Doc 

Issue 

Reason of Change Change Reference 

§5.11.2.1 New in 6.2 Additional sentence CIIS.4.1.JC.1_18 



Status 

Doc 

Issue 

Reason of Change Change Reference 

§5.1.4 6.3 TBD removed in Table CIIS.4.1.JC.2_2 

§5.1.5 6.3 Precision: Jason replaced by Jason- 

1 

§5.4.1 [PL - 5.4 -1 a] 6.3 New wording 

§5.4.1 [PL - 5.4 -2 a] 6.3 New wording PUM.6.2.EJ.15 

§5.6 [PL - 5.6 -2 ] New in 6.3 Radiation analysis requested PUM.6.2.EJ.16 

§5.11.2.2.1 Modified in 6.3 Temperature range modified PUM.6.1.EJ.27a 

§5.11.2.2.1 New in 6.3 Temperature variation added PUM.6.1.EJ.27a 

§5.11.2.2.1 Modified in 6.3 Relative humidity modified PUM.6.1.EJ.27a 





5. PAYLOAD ENVIRONMENT REQUIREMENTS 5 

5.1 MECHANICAL ENVIRONMENT 5 

5.1.1 QUASI-STATIC ACCELERATION LOADS 5 

5.1.2 SINE VIBRATION 6 

5.1.3 RANDOM VIBRATIONS 6 

5.1.4 ACOUSTICS 7 

5.1.5 PYROTECHNIC SHOCK 8 

5.2 THERMAL ENVIRONMENT 9 

5.3 DEEP SPACE VACUUM 10 

5.4 LAUNCH PRESSURE AND THERMAL FLUX PROFILES 11 

5.5 ELECTROMAGNETIC ENVIRONMENT 11 

5.6 CHARGED PARTICLES RADIATIONS 11 

5.7 MAGNETIC FIELD 15 

5.7.1 PAYLOAD SUSCEPTIBILITY 15 

5.7.2 PAYLOAD EMISSION 15 

5.8 METEROID AND SPACE DEBRIS 15 

5.9 ATOMIC OXYGEN 16 

5.10 HEAVY IONS AND TRAPPED PROTONS ENVIRONMENT 17 

5.11 GROUND OPERATIONS, STORAGE, TRANSPORTATION AND HANDLING REQUIREMENTS 

19 

5.11.1 STORAGE REQUIREMENTS 19 

5.11.2 HANDLING & TRANSPORTATION REQUIREMENTS 19 

5.11.2.1 Mechanical environment 19 

5.11.2.1.1 Road transport 19 

5.11.2.1.2 Air transport 21 

5.11.2.1.3 Handling/hoisting 21 

5.11.2.2 Thermal and climatic environment (TBC) 22 

5.11.3 INTEGRATION CONSTRAINTS 22 

5.11.4 MAINTAINABILITY 22 

5.11.5 SAFETY 22 


Figure 5.1-1 : Shock levels at payload interface ...................................................................................................... 8 

Figure 5.6-1 : Total radiation dose per year under 0.05 mm of aluminium for different inclinations ....................... 11 

Figure 5.6-2 : Total radiation dose per year under 3 mm of aluminium for different inclinations ............................ 12 

Figure 5.6-3: Radiation dose over 5 years vs Aluminium equivalent thickness and altitude ..................................... 14 

Figure 5.8-1 : Spatial Density Values in Low Earth Orbits (Jan. 1989).................................................................... 15 

Figure 5.10-1 : LET Spectrum ............................................................................................................................... 17 

Figure 5.10-2 :Trapped PROTONS Spectrum........................................................................................................ 18 

Figure 5.11-1: Shock during road transport .......................................................................................................... 20 





Table 5.1-1 : Quasi-static acceleration qualification loads (launch vehicles envelope) .............................................. 5 

Table 5.1-2 : Sine vibration input qualification levels for the payload....................................................................... 6 

Table 5.1-3 : Random vibration qualification loads ................................................................................................. 6 

Table 5.1-4 : Acoustic qualification environment ..................................................................................................... 7 

Table 5.2-1 : Sizing conditions................................................................................................................................ 9 

Table 5.2-2 : Solar constant variations.................................................................................................................... 9 

Table 5.6-1 : Radiation dose over 5 years vs Aluminium equivalent thickness and altitude ..................................... 13 

Table 5.9-1 : Material reactivity to the atomic oxygen............................................................................................ 16 

Table 5.9-2 : Annual erosion of kapton and teflon ................................................................................................ 16 

Table 5.11-1: Sine vibration during road transport................................................................................................ 19 

Table 5.11-2: Random vibration during road transport ......................................................................................... 20 

Table 5.11-3: QSL during road transport.............................................................................................................. 20 

Table 5.11-4: Sine vibration during air transport................................................................................................... 21 

Table 5.11-5: Random vibration during air transport ............................................................................................ 21 

Table 5.11-6: QSL during air transport................................................................................................................. 21 

Table 5.11-7: Acceleration during handling/hoisting............................................................................................. 21 





LIST OF TABLES...................................................................................................................................................... 3 





LIST OF TBCs 

LIST OF TBDs 

. 

N°§ Sentence Planned Resolution 

§5.11.3 TBD 




5. PAYLOAD ENVIRONMENT REQUIREMENTS 

This chapter lists the requirements about the qualification and flight environment which the equipped payload shall 

meet in order to be compatible with the PROTEUS platform. It deals with the mechanical and thermal environment, 

the deep space vacuum, the launch pressure profile, the electromagnetic, radiation, magnetic field, meteoroid and 

space debris, atomic oxygen environment. These requirements depend on the launch vehicle choice and the mission 

environment parameters (mission objective, orbit type, mission date and duration). So, as soon as these parameters 

are well defined, the User shall not hesitate to contact either ALCATEL SPACE or CNES in order to make accurate the 

payload qualification and flight environment requirements. ALCATEL SPACE and CNES can also help the User to 

define the launch vehicle and the mission parameters. 

5.1 MECHANICAL ENVIRONMENT 

The mechanical environment is caused by the launch environment. Hereafter the levels qualifying the mechanical 

environment are specified considering all the launch vehicles compatible with PROTEUS. As soon as the considered 

launch vehicles envelope is restrained (because some or one launch vehicle is chosen among the specified launch 

vehicles for the studied mission), the mechanical levels are reduced. Therefore, the payload environment is less 

constrained, the payload design requirements are less severe. 

5.1.1 QUASI-STATIC ACCELERATION LOADS 

PL - 5.1.1 -1 

The quasi static qualification load factors for the payload are given in Table 5.1-1, in satellite axes. 

For information, these quasi static loads are not applicable to the secondary structures or instruments because they 

are covered by dynamic vibration loads. 

Longitudinal Qualif. Load (g) Lateral Qualif. Load (g) 

Payload 20 9 - 0,02 x (M-100) for 100 kg < M < 200 kg 

7 - 0,005 x (M-200) for 200 kg < M < 300 kg 

Table 5.1-1 : Quasi-static acceleration qualification loads (launch vehicles envelope) 

Lateral and longitudinal QS loads have not to be combined. 

There is no quasi static test requirement since payload strength will be tested during sine vibration testing. The quasi 

static qualification load factors given in this section shall only be used to structurally design the payload. 




5.1.2 SINE VIBRATION 

PL - 5.1.2 -1 

The payload shall withstand the sine vibration input qualification levels given in Table 5.1-2, in satellite axes. 

These preliminary values are taken from PROTEUS (specified launch vehicles envelope) and will be refined after 

coupled platform/payload mechanical analysis and launch vehicle inputs. 

Part Excitation Axis Frequency Range Input Level (QL) 

Payload 

longitudinal 

(Xs) 

lateral 

(Ys, Zs) 

5 -> 21 Hz 

21 -> 30 Hz 

30 -> 50 Hz 

50 -> 100 Hz 

5 -> 14 Hz 

14 -> 20 Hz 

20 -> 40 Hz 

40 -> 80 Hz 

80 -> 100 Hz 

11 mm 

20 g 

linear connection 

5 g 

11 mm 

9 g 

5 g 

1.5 g 

3 g 

Table 5.1-2 : Sine vibration input qualification levels for the payload 

The sine levels in the low frequency range (


5.1.4 ACOUSTICS 

PL - 5.1.4 -1 

The payload shall withstand the qualification sound pressure levels defined Table 5.1-4. 

These levels are the envelope of 4 launch vehicles (Rockot, PSLV, Delta 2, Soyuz) for which PROTEUS is compatible. 

As soon as the mission specific launch vehicle is chosen, these levels will be updated and recorded in the Payload 

Design Interface Specification 

Octave Band Center Frequency 

(Hz) 

Qualification levels 

(dB) 

31.25 128.5 

62.5 135 

125 137 

250 140 

500 141 

1000 136 

2000 132 

4000 129 

8000 126 

Overall 146 

Table 5.1-4 : Acoustic qualification environment 

The acceptance levels are 3 dB lower than qualification levels. 




5.1.5 PYROTECHNIC SHOCK 

PL - 5.1.5 -1 

The following shock levels are applicable on the payload at the mating interface (at each pod upper face). 

The shock levels experienced by the payload comes from the launch vehicle separation and from the solar 

arrays deployment. 

shock level (g) 

10000 

1500 

1000 

100 

Q = 10 

10 

100 1000 


2000 

10000 

Figure 5.1-1 : Shock levels at payload interface 

As explained in PL-5.1.5-1 requirement, this spectrum is given at the PF/PL interface plane level. The overall payload 

can be tested with this level, but it is generally difficult to perform such a test. 

In case of verification at payload equipment level only (test philosophy to be analysed by CNES on the basis of the 

payload validation plan delivered by the Payload Supplier), it may be noticed that, in the framework of the JASON-1 

program, the payload equipment had been successfully qualified with a shock spectrum corresponding to an half 

sine of 900 g amplitude and 0.5 ms duration. This payload equipment level qualification has allowed to cover shock 

levels measured during JASON-1 satellite shock tests, even for the equipments very close to the PF/PL interface. 

It may be noticed that the payload shall not generate shock levels higher than those required in section 3.1.5.2 (for 

PF/PL interface plane) and Section 3.6.2.3.3 (for PL/STA interface plane). 




5.2 THERMAL ENVIRONMENT 

The payload is submitted to albedo, Earth and Sun fluxes. 

PL - 5.2 -1 

Sizing conditions (orbital parameters, satellite attitude) are given in Table 5.2-1. 

SATELLITE MODE DURATION APPARENT DIRECTION OF 

THE SUN 

Launch phase (payload thermal 

control OFF) 

Mission dependent (up to 90 

min) 

Mission dependent 

SHM mode (RDP and SPP phases) 360 min Random 

SHM mode (BBQ phase) Unlimited -X s ± 30° 

Normal mode Mission dependent 

Transient Mission dependent 

Table 5.2-1 : Sizing conditions 

The following data will be incorporated in mission environment specifications when it is issued. 

PL - 5.2 -2 

• solar constant : the solar constant variations are given in Table 5.2-2. The ± 5 W/m2 variation is 

due to the 11 year solar cycle. 

PL - 5.2 -3 

TIME OF YEAR SOLAR CONSTANT (W/m 2 ) 

Winter solstice (perihelion) 1415 ±5 

Vernal equinox 1380 ±5 

Summer solstice (aphelion) 1326 ±5 

Autumnal equinox 1365 ±5 

Table 5.2-2 : Solar constant variations 

• albedo and Earth infrared (IR) fluxes: the Earth and its atmosphere radiate like a black body at an 

equivalent temperature of 255 K. The albedo coefficient is the ratio of the Earth reflected solar flux by the 

overall incident solar flux. The mean value of the albedo coefficient is 0.3 but this value varies from zone 

to zone on Earth. The same albedo coefficient shall be considered for the albedo and Earth IR fluxes 

calculations. This yields the following formulas: 

• albedo flux = albedo coefficient x solar flux 

• Earth IR flux = (1 - albedo coefficient)/4 x solar flux 

• with albedo coefficient = 0.3 ±0.05 




PL - 5.2 -4 


• deep space flux: the deep space thermal radiation is equivalent to a black body at a 4 K 

temperature. 

5.3 DEEP SPACE VACUUM 

PL - 5.3 -1 

The payload shall withstand the deep space vacuum conditions. Free space vacuum pressure to be 

considered in orbit life is below 10 -8 Pa. 




5.4 LAUNCH PRESSURE AND THERMAL FLUX PROFILES 

PL - 5.4 -1 a 

The payload shall withstand an expected maximum pressure decay during the launch ascent phase up to 

4000 Pa/s. 

PL - 5.4 -2 a 

The payload shall withstand the aerothermal flux after fairing jettisoning lower than 1135 W/m 2 . 

5.5 ELECTROMAGNETIC ENVIRONMENT 

PL - 5.5 -1 

The design shall comply with requirements of Section 3.5.7 regarding design guidelines and of section 6.1.8 

regarding test procedures and set-up. 

5.6 CHARGED PARTICLES RADIATIONS 

The dose of radiation received by the payload depends on the satellite orbit. The yearly received doses depending on 

the altitude and the inclination of the orbit are shown for two typical equivalent of aluminium thicknesses : 

.. 

0.05 mm which corresponds to an external dose (cf. Figure 5.6-1), 

3.0 mm which corresponds to a minimal shielding : 2 mm brought by the structure, 1 mm brought by the 

studied equipment box and the neighbouring equipment (cf. Figure 5.6-2) 

Figure 5.6-1 : Total radiation dose per year under 0.05 mm of aluminium for different 

inclinations 




Figure 5.6-2 : Total radiation dose per year under 3 mm of aluminium for different 

inclinations 

The radiation dose received at EEE parts level is a function of the protection given by: 

the other units and the satellite structure, 

the unit box and the other elements of the unit. 

PL - 5.6 -1 

The payload sizing shall take into account the total radiation dose (4 pi steradian) versus shielding protection 

thickness (margins included). 

PL - 5.6 -2 

The Payload Supplier shall provide the Satellite Supplier with a radiations analysis at parts level, accounting 

for every protection (satellite structure, unit structure, other electronics parts), and with the radiation dose 

versus thickness given hereafter. 

Table 5.6-1 presents, for information, the maximal radiation dose cumulated over 5 years for different Aluminium 

equivalent thickness shielding and for different altitudes. 

Table 5.6-1 presents, for information, the maximal radiation dose cumulated over 5 years for different Aluminium 

equivalent thickness shielding and for different altitudes. 

Figure 5.6-3 gives a graphical representation of the data provided in Table 5.6-1. 



.. 


Aluminium Equivalent thickness Radiation dose (rad) 

(mm) 700 km 900 km 1100 km 1336 km 

Jason orbit 



1336 km 

Worst case 

0 7.65E+06 1.31E+07 3.38E+07 2,96E+07 8.40E+07 

0.1 1.30E+06 4.47E+06 1.22E+07 1,85E+07 3.11E+07 

0.2 5.90E+05 1.84E+06 5.12E+06 8,24E+06 1.32E+07 

0.3 3.17E+05 7.53E+05 2.15E+06 4,11E+06 5.53E+06 

0.4 2.00E+05 3.87E+05 1.11E+06 2,22E+06 2.87E+06 

0.5 1.38E+05 2.26E+05 6.50E+05 1,32E+06 1.69E+06 

0.6 1.02E+05 1.48E+05 4.23E+05 8,67E+05 1.10E+06 

0.8 6.39E+04 7.74E+04 2.19E+05 4,71E+05 5.74E+05 

1 4.52E+04 5.00E+04 1.39E+05 3,11E+05 3.65E+05 

1.5 2.54E+04 2.65E+04 7.04E+04 1,60E+05 1.81E+05 

2 1.60E+04 1.87E+04 4.78E+04 1,01E+05 1.20E+05 

2.5 1.08E+04 1.51E+04 3.75E+04 7,03E+04 9.21E+04 

3 7.52E+03 1.29E+04 3.15E+04 5,17E+04 7.58E+04 

4 4.15E+03 1.04E+04 2.45E+04 3,20E+04 5.77E+04 

5 2.58E+03 8.96E+03 2.06E+04 2,25E+04 4.75E+04 

6 1.87E+03 8.15E+03 1.84E+04 1,82E+04 4.23E+04 

7 1.54E+03 7.67E+03 1.72E+04 1,60E+04 3.92E+04 

8 1.37E+03 7.31E+03 1.63E+04 1,49E+04 3.69E+04 

9 1.27E+03 6.94E+03 1.56E+04 1,39E+04 3.51E+04 

10 1.21E+03 6.69E+03 1.50E+04 1,31E+04 3.35E+04 

12 1.12E+03 6.23E+03 1.39E+04 1,19E+04 3.11E+04 

14 1.04E+03 5.88E+03 1.30E+04 1,11E+04 2.91E+04 

16 9.76E+02 5.53E+03 1.22E+04 1,03E+04 2.73E+04 

18 9.15E+02 5.22E+03 1.16E+04 9,56E+03 2.58E+04 

20 8.69E+02 4.96E+03 1.10E+04 9,19E+03 2.45E+04 

Table 5.6-1 : Radiation dose over 5 years vs Aluminium equivalent thickness and altitude


Radiation dose over 5 years (rad) 

1,00E+06 

1,00E+05 

1,00E+04 

1,00E+03 

700 km 

900 km 

1100 km 

1,00E+02 

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 

Aluminium Equivalent Thickness (mm) 



1336 km 

(JASON1 orbit) 

1336 km 

(JASON1 orbit worst case) 

Figure 5.6-3: Radiation dose over 5 years vs Aluminium equivalent thickness and altitude


5.7 MAGNETIC FIELD 

5.7.1 PAYLOAD SUSCEPTIBILITY 

PL - 5.7.1 -1 

Deleted (see PL - 3.5.9 - 2). 

PL - 5.7.1 -2 


5.7.2 PAYLOAD EMISSION 

PL - 5.7.2 -1 

Deleted (See PL - 3.5.9 - 1). 

PL - 5.7.2 -2 


5.8 METEROID AND SPACE DEBRIS 

For information, Figure 5.8-1 plots spatial density values out to 2000 km altitude for trackable objects of various 

sizes. 

Figure 5.8-1 : Spatial Density Values in Low Earth Orbits (Jan. 1989) 




5.9 ATOMIC OXYGEN 

PL - 5.9 -1 

The payload may be exposed to the atomic oxygen environment. The design shall consider performance in 

this environment. 

The atomic oxygen, mostly fixed, follow the rotation of the Earth and its atmosphere. Therefore, they hit the satellite 

front face with a velocity around 26000 km/h. This kinetic energy adds to the high chemical reactivity of oxygen 

atoms that imply a fast reaction with the hit materials. A chemical effect occurs and induces a materials surfaces 

fragilization and a mechanical effect with a materials surfaces erosion. The erosion thickness depends on the oxygen 

dose and the materials kind. Table 5.9-1 gives the reactivity of main materials usually used in space technology. 

Material Erosion 10-24 cm3 /atom 

Kapton H polyimide 3.0 

Mylar polyester 2.7 to 3.9 

Polyethylene 3.3 to 3.7 

Epoxy 1.7 

Polycarbonate 2.9 to 6.0 

Polystyrene 1.7 

Polysulfone 2.4 

Urethane (black, conductor) 0.3 

Silver 10.5 

Carbon 0.9 to 1.7 

Chemglaze Z306 (cblack) paint 0.35 

FEP Teflon 0.037 to 0.35 

Aluminium 0.0 

Copper 0.0 

Gold 0.0 

SiO2 0.0 

Table 5.9-1 : Material reactivity to the atomic oxygen 

Table 5.9-2 gives some typical values of eroded thickness per year for a mean solar activity. 

Altitude 

Oxygen flux kapton erosion (µm) Teflon erosion (µm) 

(km) 

(atomes/cm2.s) 

300 8.1014 750 87 

400 1014 95 10 

500 2.1013 20 2.3 

600 4.1012 3.8 0.45 

700 1012 0.95 0.1 

800 2.1011 0.20 0.023 

900 4.1010 0.04 0.005 

1000 8.109 0.008 0.001 

Table 5.9-2 : Annual erosion of kapton and teflon 




5.10 HEAVY IONS AND TRAPPED PROTONS ENVIRONMENT 

PL - 5.10 -1 

The payload may be exposed to the heavy ions and trapped protons environment. The design shall consider 

performance in this environment. 

The orbital environment in terms of LET (Linear Energy Transfer) and Trapped Protons for a worst case in the 

PROTEUS flight domain (z = 1336 km) is shown on Figure 5.10-1 and Figure 5.10-2 (corresponding to the Jason 

case). It includes contributions from solar flares. 

.. 

Flux (Part m -2 ster -1 s -1 ) 

1.E+04 

1.E+03 

1.E+02 

1.E+01 

1.E+00 

1.E-01 

1.E-02 

1.E-03 

1.E-04 

1.E-05 

1.E-06 

1.E-07 

1.E-08 

1.E-09 

1.E-10 

1.E-11 

0 20 40 60 80 100 120 

Linear Energy Transfer (MeV mg -1 cm 2 ) 

Figure 5.10-1 : LET Spectrum 




Flux (cm -2 jour -1 ) 

8.E+07 

7.E+07 

6.E+07 

5.E+07 

4.E+07 

3.E+07 

2.E+07 

1.E+07 

0.E+00 

0 50 100 

Energy (MeV) 

150 200 

Figure 5.10-2 :Trapped PROTONS Spectrum 




5.11 GROUND OPERATIONS, STORAGE, TRANSPORTATION AND HANDLING REQUIREMENTS 

5.11.1 STORAGE REQUIREMENTS 

PL - 5.11.1 -1 

Any payload shall be able to withstand a storage period of 6 months after payload delivery to the satellite 

added to 1,5 year between the AIT and the launch without degradation of its functions or performance 

This storage will occur under the following conditions : 

temperature : 20° C ± 10° C 

relative humidity : 40% ± 20% 

5.11.2 HANDLING & TRANSPORTATION REQUIREMENTS 

PL - 5.11.2 -1 

It shall be possible to transport the payload integrated on the satellite with the environment described in the 

following paragraphs 

5.11.2.1 Mechanical environment 

The static and dynamic mechanical environment affecting the payload during all ground operations is covered by the 

envelope of the defined launch mechanical environment (i.e. ground operations shall not drive the design). 

The mechanical environment is generated by air/road transportation and handling. 

Factors of safety are given in section 4.2.5.2. 

Following acting loads are transportation/handling loads. 

5.11.2.1.1 Road transport 

Sine vibration 

The following accelerations act in any three axes simultaneously 

Frequency range Level 

10 Hz - 100 Hz 1.2 g 

Table 5.11-1: Sine vibration during road transport 

3-axis random vibration (standard environment) 




Frequency range Level Global 

(g RMS) 

5 Hz - 10 Hz + 6 dB/oct. 

10 Hz - 100 Hz 0.003 g²/Hz 0.64 

100 Hz - 200 Hz -12 dB/oct 

200 Hz - 400 Hz 0.0001875 g²/Hz 

Table 5.11-2: Random vibration during road transport 

Shock 

10 g during 10 ms according to the following shock profile. 

Acceleration (g) 

20 

10 

Quasi-static (40 km/h top speed) 

0 

0 5 10 15 20 

Duration (ms) 

Figure 5.11-1: Shock during road transport 

X Y Z 

± 1.1 g ± 1.2 g +1.0 g / -3.0 g 

* X velocity, Z vertical 

Table 5.11-3: QSL during road transport 

Accelerations act simultaneously along all the 3 axes. 




5.11.2.1.2 Air transport 

3-axes sine vibration (standard environment) 

Frequency band Level 

2- 20 Hz ± 0.2 mm 

20- 50 Hz 0.85 g 

50-100 Hz 2 g 

Table 5.11-4: Sine vibration during air transport 

3-axis random vibration (standard environment) 

Frequency range Level Global (g RMS) 

5 Hz - 10 Hz + 6 dB/oct. 

10 Hz - 100 Hz 0.003 g²/Hz 0.64 

100 Hz - 200 Hz -12 dB/oct 

200 Hz - 400 Hz 0.0001875 g²/Hz 

Table 5.11-5: Random vibration during air transport 

Shock : 

Half sine profile of 4.2 g amplitude and 20 ms duration 

Quasi-static 

Aircraft axis X (forward) Y Z (+ up) 

Landing + 1.5 g ± 1.5 g -2.0 g 

Take-off - 1.5 g 0 g +2.0 g / -1.5 g 

Table 5.11-6: QSL during air transport 

Accelerations act simultaneously along all the 3 axes. 

5.11.2.1.3 Handling/hoisting 

Acceleration 

Hoisting sling VERTICAL 

HORIZONTAL 

1.3 g 

0.1 g } Act simultaneously 

Assembly and integration jig VERTICAL 1.5 g 

} Act simultaneously 

Table 5.11-7: Acceleration during handling/hoisting 

The gravity acceleration is included in the vertical acceleration 




Shock 


Equivalent to a 10 cm drop, considering that there is already a MGSE / ground contacting point. 

5.11.2.2 Thermal and climatic environment (TBC) 

The thermal and climatic environment during transportation is : 

Temperature : in the [5°C, 50°C] range 

Temperature variation : +5° C/h maximum 

Relative humidity : < 55 % 

Cleanliness : better than class 100000 

5.11.3 INTEGRATION CONSTRAINTS 

TBD 

5.11.4 MAINTAINABILITY 

PL - 5.11.4 -1 

The payload shall be designed to require a minimum of special tools and test equipment to maintain 

calibration, perform adjustments and accomplish fault identification 

Marking and location of the connectors shall be easily distinguished without any mistake possibility 

5.11.5 SAFETY 

PL - 5.11.5 -1 

The payload shall comply with the standard rules for the utilisation in Clean Room. 





Chapter 6 : Payload verification and test requirements 


Here below are listed the changes between issue N-2 and the issue N-1: 


Status 


Issue 

§6.1.2.1 New in Aims of the validation tests PUM.6.1.EJ.29 6.2 

§6.1.5.1 New in § at a different level: Electrical 

Functional Verification 



PUM.6.1.CG.31_25 6.2 

§6.1.5.1 [PL - 6.1.5 -4 ] New in Kinds of payload reference test PUM.6.1.CG.31_25 6.2 

§6.1.5.1 [PL - 6.1.5 -5 ] New in Duration of the tets PUM.6.1.CG.31_25 6.2 

§6.1.5.2 New in § Electrical Interface Validation PUM.6.1.CG.31_25 6.2 

§6.1.6.1 [PL - 6.1.6 -3 a] Modified in Table replace by Section PUM.6.1.CG.31_26 6.2 

§6.1.6.4 [PL - 6.1.6 -8 a] Modified in Handling: additional sentence PUM.6.1.CG.31_27 6.2 

§6.2 Modified in Figure 6.2-1 updated PUM.6.1.CG.31_28 6.2 

§6.2.5.1 New in § 6.2.5.1 updated PUM.6.1.CG.31_30 6.2 

§6.2.5.2 Modified in § 6.2.5.2 updated PUM.6.1.CG.31_30 6.2 



Status 

§6.1.5.2.1 Deleted 

in 

§6.1.5.2.1 [PL - 6.1.5 -7 ] Modified 

in 


Issue 

Sentence deleted PUM.6.1.CG.31_25a 6.3 

One sentence added PUM.6.1.CG.31_25a 6.3 

§6.1.6.3.4.4 [PL - 6.1.8 -29 ] New in Text becomes a requirement PUM.6.2.E.J.21 6.3 


in 


in 

Figure updated PUM.6.1.CG.31_30a 6.3 

RF launcher required bands ranges 

specified in section 3.5.7.2.1 

PUM.6.2.EJ.33 6.3 

§6.2.5.2 Modified RF Susceptibility defined in Section PUM.6.2.EJ.33 6.3


N°§ PUID Change Reason of Change Change Reference Doc 

Status 

Issue 

in 3.5.7.2.2 

§6.2.5.2 Deleted 

in 

Launcher TM band deleted PUM.6.2.EJ.33 6.3 


CHANGE TRACEABILITY CHAPTER 6 1 

6. CHAPTER 6: PAYLOAD VERIFICATION AND TEST REQUIREMENTS 7 

6.1 PAYLOAD DESIGN VERIFICATION REQUIREMENTS 7 

6.1.1 GENERAL 7 

6.1.2 PAYLOAD MODEL BUILD STANDARD 7 

6.1.2.1 Payload Functional Model definition (TBC) 8 

6.1.2.2 Qualification and Flight Spares (QFS) definition 8 

6.1.2.3 ProtoFlight Model (PFM) definition 8 

6.1.2.4 Flight Model (FM) definition 8 

6.1.3 DESIGN VERIFICATION METHODS AND TYPES - DEFINITION 9 

6.1.3.1 Verification Methods 9 

6.1.3.1.1 Functional Tests 9 

6.1.3.1.2 Environmental Tests 9 

6.1.3.1.3 Verification by Similarity 9 

6.1.3.1.4 Verification by Analysis 9 

6.1.3.1.5 Verification by Inspection 9 

6.1.3.1.6 Verification by Demonstration 9 

6.1.3.1.7 Verification by Validation of Records 10 

6.1.3.2 Verification Types 10 

6.1.3.2.1 Development Verification 10 

6.1.3.2.2 Qualification Verification 10 

6.1.3.2.3 Acceptance Verification 10 

6.1.4 GENERAL REQUIREMENTS FOR MEASUREMENTS AND TESTS 11 

6.1.4.1 Environmental Conditions 11 

6.1.4.2 Tolerance levels 11 

6.1.4.3 Cleanliness of test equipment 13 

6.1.4.4 Measurements 13 

6.1.5 ELECTRICAL VERIFICATION 14 

6.1.5.1 Electrical Functional Verification 14 

6.1.5.1.1 Payload Performance Verification Test (PPVT) 14 

6.1.5.1.2 Payload Health Check Test (PHCT) 14 

6.1.5.1.3 Payload Aliveness Test (PAT) 14 

6.1.5.2 Electrical Interfaces Validation 15 

6.1.5.2.1 Platform interfaces simulation 15 

6.1.5.2.2 Electrical interface validation 15 

6.1.5.2.2.1 Pin allocation/signal addressing 15 

6.1.5.2.2.2 Continuity/Isolation 16 

6.1.5.2.2.3 Primary Power lines 16 

6.1.5.2.2.4 Heater lines 16 

6.1.5.2.2.5 Other lines 16 




6.1.6 STRUCTURAL AND MECHANICAL VERIFICATION 18 

6.1.6.1 Sinusoidal Vibrations 18 

6.1.6.2 Acoustic/Random Vibrations 19 

6.1.6.3 Shocks 20 

6.1.6.4 Handling 20 

6.1.7 THERMAL VERIFICATION 21 

6.1.7.1 Thermal Balance Test definition 21 

6.1.7.2 Thermal Vacuum Test (Thermal Cycling) definition 21 

6.1.8 EMC VERIFICATION 22 

6.1.8.1 Test Configuration 22 

6.1.8.2 Test Requirements 23 

6.1.8.2.1 Units and harness/wiring configuration 23 

6.1.8.2.2 Test operating conditions 23 

6.1.8.2.3 Band analyses 24 

6.1.8.2.4 Amplitude 24 

6.1.8.3 Responsibilities 24 

6.1.8.4 Test Site 25 

6.1.8.4.1 Facility requirements 25 

6.1.8.4.2 Tests outside an anechoic chamber 25 

6.1.8.4.3 Measuring instrument 26 

6.1.8.4.3.1 Measuring receiver 26 

6.1.8.4.3.2 Spectrum analyser 26 

6.1.8.4.3.3 Electric field measurement antennas 26 

6.1.8.4.3.4 Calibration 26 

6.1.8.4.4 Test set-ups 27 

6.1.8.5 TESTS 28 

6.1.8.5.1 Conducted test requirements 28 

6.1.8.5.2 Radiated test requirements 30 

6.1.8.5.3 Electrical Ground Support Equipment (EGSE) 30 

6.1.8.6 Tests organization 31 

6.1.8.6.1 Test Plan 31 

6.1.8.6.2 Test procedure 31 

6.1.8.6.3 Test execution 32 

6.1.8.6.4 Presentation of results 32 

6.1.8.6.4.1 General 32 

6.1.8.6.4.2 Test report 33 

6.1.8.7 Unit Test set-ups 34 

6.1.8.7.1 Conducted emissions; Power supply lines, steady perturbations 34 

6.1.8.7.2 Conducted emissions; Power supply lines, transient perturbations 35 

6.1.8.7.3 Conducted susceptibility ; power supply lines, sine wave and square wave 36 

6.1.8.7.4 Conducted Susceptibility; power supply lines, transient signal 37 

6.1.8.7.5 Susceptibility to common mode transients; interface signals 38 

6.1.8.7.6 Radiated emissions E-fields 39 

6.1.8.7.7 Radiated susceptibilities E-fields 40 

6.1.8.7.8 Magnetic moment (DC) 40 

6.1.9 ESD VERIFICATION 41 

6.1.10 MAGNETIC FIELD VERIFICATION 41 

6.1.11 VERIFICATIONS PRIOR TO PAYLOAD DELIVERY 42 

6.1.11.1 Inspections and examinations at unit level 42 

6.1.11.2 Mass properties determination 42 

6.1.11.3 Unit acceptance and delivery for satellite integration 42 

6.2 TESTS AND VERIFICATIONS AT SATELLITE LEVEL 43 

6.2.1 PAYLOAD INSPECTION BEFORE INTEGRATION 44 

6.2.2 FUNCTIONAL TESTS 44 




6.2.3 THERMAL VACUUM TESTS 45 

6.2.4 VIBRATION TESTS 47 

6.2.4.1 Sinusoidal Vibrations 47 

6.2.4.2 Random Vibrations 47 

6.2.4.3 Acoustic Noise 47 

6.2.4.4 Pyrotechnic shocks 47 

6.2.5 EMC-TEST 48 

6.2.5.1 Conducted emission 48 

6.2.5.2 Radiated emission and susceptibility 48 

6.2.5.3 RF compatibility test 50 

6.2.6 ESD-TEST 50 


Figure 6.1-1: Schematic representation of a LISN.................................................................................................. 28 

Figure 6.1-3 : Power lines, steady perturbations test set up.................................................................................... 34 

Figure 6.1-4 : Power supply line, transient perturbations test set-up ....................................................................... 35 

Figure 6.1-5 : Conducted susceptibility test set-up (sine wave and square wave) .................................................... 36 

Figure 6.1-6 : Conducted susceptibility test set-up (transient signal) ....................................................................... 37 

Figure 6.1-7 : Interface signals test set-up............................................................................................................. 38 

Figure 6.1-8 : Radiated emissions E-fields test set-up............................................................................................. 39 

Figure 6.1-9 : Radiated susceptibilities E-fields test set-up...................................................................................... 40 

Figure 6.2-1 : Satellite Assembly Integration and Test............................................................................................ 43 


Table 6.1-1 : Tests required at payload level before payload delivery ...................................................................... 7 

Table 6.1-2 : Measurement requirements.............................................................................................................. 12 

Table 6.1-2b: Payload electrical interface verification matrix.................................................................................. 17 

Table 6.1-3 : Acoustic test tolerances .................................................................................................................... 19 

Table 6.1-4: Analysis bandwidth for NB and BB recordings ................................................................................... 24 

Table 6.1-5: Measuring range and section for CE and CS tests ............................................................................. 29 

Table 6.1-6: Measuring range and section for RE and RS tests............................................................................... 30 

Table 6.2-1: Functional tests................................................................................................................................. 44 









LIST OF TBCs 

List of TBDs 

§ N° Sentence Planned Resolution 

§6.2.5.2 S/C EMC Radiated Emission and Susceptibility with launcher TBD: 

§6.2.5.3 Test sequence TBD. 




Chapter 6: Payload verification and test requirements 

The first section details the "rules of the art" and provides preliminary requirements to qualify the payload before 

delivery and consequently ensure the best compatibility with the satellite (especially for EMC). The second section 

describes the main tests and verifications (instrument inspection, functional tests) at satellite level in order to give a 

rough idea of the satellite tests sequence for the User. 

6.1 PAYLOAD DESIGN VERIFICATION REQUIREMENTS 

6.1.1 General 

PL - 6.1.1 -1 

The demonstration of the qualification status shall be given to the Satellite Contractor through the 

Development, Design and Verification (DD&V) documents and through the Payload End Item Data Package 

as defined in the Deliverable Items List. 

PL - 6.1.1 -2 

The Payload shall be delivered to the Satellite Contractor fully qualified. Required tests are given in Table 

6.1-1. 

Kind of tests Required Comments 

PVT X For reference test purpose. 

Functional HCT X For reference test purpose. 

AT X For reference test purpose. 

Sine X 

Mechanical Acoustic or random X Depending on payload shape 

Shock X At payload or payload sub-system level 

Thermal Thermal balance X 

Thermal cycling X 

CE/CS X Test set-up described in section 6.1.8.7 

EMC RE/RS X Test set-up described in section 6.1.8.7 

ESD ESD X 

Mass properties Mass properties X Mass and inertia 

Table 6.1-1 : Tests required at payload level before payload delivery 

The payload development and verification philosophy shall contain, at least, these required tests. Full test campaign 

and test levels and duration (qualification and/or acceptance) shall be determined by the Payload Supplier 

depending on the payload maturity and agreed by the Satellite Contractor. 

The Payload Supplier shall deliver to the Satellite Supplier the Payload Flight Model and the Spare Model (if 

applicable). 

6.1.2 Payload Model Build Standard 

The payload level of assembly and build standard shall comply with the System verification concept selected for the 

PROTEUS based mission Program. 

Provisions for payload models have to be made according to the Deliverable Items List. 




PL - 6.1.2 -1 

The payload model philosophy and its related payload qualification/acceptance program shall be defined by 

the Payload Supplier in his Payload Development and Validation Plan. 

Hereafter are defined the payload models usually assembled and built to check the payload concept and 

performances. With this information, the User can estimate the need to build or not all these models according to his 

mission. 

6.1.2.1 Payload Functional Model definition (TBC) 

The Payload Functional Model shall be representative of the Flight Model for the following aspects: 

Electrical interface parameters 

All interface hardware shall be electrically and functionally representative of the flight standard (excluding use 

of high reliability parts). 

Command control interface 

All interface hardware and software equipment shall be representative of the flight standard. 

Connectors interface 

The interface connectors shall be flight representative. In the event the connectors interface with flight 

hardware or EGSE that also interfaces with flight hardware, gold plated hi-rel type or connector savers shall be 

used. 

This model will be used for functional tests at satellite level on satellite validation bench. Requirements for this model 

are given in document reference «PIC-P0.3-NT-224-CNES». 

Main aims of these validation tests are: 

Test of the PF-PL communications for the mission: 

• 1553 dialog: 

1553 TCs: Payload Controller Commands, Payload Software 

PLTM 

broadcast command: PPS UTC date message 

• discrete acquisition lines from the Payload (OBDH addressing) 

• discrete commands from the Platform to the Payload (OBDH addressing) 

• Payload software loading through the Platform 

FDIR testing 

• For example: - Following to 3 consecutive out of range current values acquisitions on line n°X of the 

Payload, opening by the Platform of the Payload power lines relays according to a predefined order. 

• Following to 3 consecutive out of range PLC temperature value acquisitions, opening by the Platform of the 

Payload power lines relays according to a predefined order. 

Payload interface level tests 

• For example: - closure of the power lines relays according to a predefined order with a fixed timing 

• - discrete command sensivity and observation of PL status change 

System level tests 

• All the functional chains together, with the modes chaining simulation according to real time performances. 

6.1.2.2 Qualification and Flight Spares (QFS) definition 

The objective is to qualify Payload off-line of the system qualification program. 

This Payload shall be used as flight spare. Refurbishment of Payloads shall therefore be considered. 

6.1.2.3 ProtoFlight Model (PFM) definition 

The ProtoFlight Model shall be of a standard compliant with all the requirements of the applicable Payload Design 

Interface Specification (PDIS) last issue and shall have successfully undergone a full program of qualification testing 

(with acceptance duration) and verification prior to delivery. 

6.1.2.4 Flight Model (FM) definition 

The Flight Model shall be of a standard compliant with all the requirements of the applicable PDIS last issue and 

shall have successfully undergone a full program of acceptance testing and verification prior to delivery. 




6.1.3 Design Verification Methods and Types - Definition 

The Payload Development and Validation plan must be based on a development qualification and acceptance 

scheme compatible with the overall system program concept. 

6.1.3.1 Verification Methods 

Qualification and Acceptance Verification shall be accomplished by test wherever possible and by assessment as 

support or as an alternative should testing be prohibitive: 

a) Test 

.Functional Tests, 

. Environmental Tests, 

b) Assessment 

. Similarity, 

. Analysis, 

.Inspection, 

. Demonstration, 

. Validation of Records. 

6.1.3.1.1 Functional Tests 

Functional testing is a series of electrical or mechanical performance tests conducted on flight or flight configured 

hardware at conditions equal or less than design specifications. Its purpose is to establish that the hardware performs 

satisfactorily in accordance with the design specifications. Depending on the situation, there are functional tests of 

various complication or degrees of depth. 

6.1.3.1.2 Environmental Tests 

An environmental test is a test conducted on flight or flight configured hardware to assure that the flight hardware 

will perform satisfactorily in one or more of its flight environments. Example are acoustic, thermal vacuum and EMC. 

Environmental testing is normally combined with functional testing to a degree which depends on the objectives of 

the test. 

6.1.3.1.3 Verification by Similarity 

Verification by similarity is the process of assessing by review that the article is similar or identical in design and 

manufacture to another article that has previously been qualified to equivalent or more stringent conditions. 

6.1.3.1.4 Verification by Analysis 

Verification by analysis is a process where compliance of an article to specification is proven by analytical methods. 

The typical technique used is mathematical modeling (e.g. by finite elements method, simulation, statistics, etc.). 

Mathematical models may be supplemented or supported by hardware simulations. Verification by analysis is 

normally given lower importance than direct testing, but is applicable where: 

Analysis is rigorous and accurate enough to provide reliable results, 

Tests are not cost effective, 

Similarity is not available. 

6.1.3.1.5 Verification by Inspection 

Verification by inspection may typically be applied where an article consists of well known and proven manufacturing 

methods. The verification process consists in assuring strict adherence to these specified methods during the article 

production (i.e. exclusion of deviations and mistakes) by rigorous supervision and inspection (e.g. wire coding, 

materials selection, mechanical and electrical connections, correct screw torquing, etc.). Depending on the specific 

case, inspection may reduce or omit later testing of the article in various aspects. Like analysis, inspection in general 

is given lower priority than direct testing, but may be applied where other verification methods are not cost effective. 

6.1.3.1.6 Verification by Demonstration 

Verification by demonstration primarily applies to activities of a handling, servicing, safety and logistics nature, e.g. 

easy replaceability of a critical Payload unit, lifting a container with fork-lift, mounting a Payload on a vibrator, etc. 




The process consists in demonstrating that the activity in question is possible within the specified time, manpower, 

safety and other constraints. 

6.1.3.1.7 Verification by Validation of Records 

Verification by validation of records is a process where on the basis of manufacturing records (which have to be 

complete and comprehensive, and may not contain any new unproved processes), compliance with performance 

specifications of an article can be proven. This process is of the same nature as inspection, it being an inspection of 

(reliable) records "after the event". Again, this verification method is considered lower priority and applies if direct 

testing is not feasible. 

6.1.3.2 Verification Types 

6.1.3.2.1 Development Verification 

Development Verification is a process to verify the feasibility of a design approach and to provide confidence in the 

ability of the hardware to comply with the performance criteria. 

6.1.3.2.2 Qualification Verification 

Qualification Verification is an individual test or a series of functional and environmental tests conducted on flight 

hardware at conditions normally more severe than acceptance test conditions, to establish that the hardware will 

perform satisfactorily in the flight environments with sufficient margins. The purpose is to uncover deficiencies in 

design and method of manufacture. It is not intended to exceed design safety margins or to introduce unrealistic 

modes of failure. 

6.1.3.2.3 Acceptance Verification 

Acceptance Verification is an individual test or a series of functional and environmental tests conducted on flight 

hardware at conditions equal to design specifications plus acceptance level margin to establish that the hardware 

performs satisfactorily. 




6.1.4 General Requirements for measurements and tests 

6.1.4.1 Environmental Conditions 

PL - 6.1.4 -1 

All measurements and tests shall be conducted within the following environmental conditions: 

• Pressure: ambient 

• Temperature: 22°C + 3°C 

• Relative Humidity: < 60 %. 

PL - 6.1.4 -2 

Actual ambient test conditions shall be recorded regularly during the tests. In case of ambient conditions 

exceeding the allowable limits, the decision not to test or to halt any test in progress shall lie with the 

responsible Test Manager who must have adequate evidence that there will be no adverse influences on 

component performance. 

6.1.4.2 Tolerance levels 

PL - 6.1.4 -3 

The accuracy of instruments and test equipment used to control or monitor the test parameters shall be better 

than one tenth of the tolerance of the variable to be measured. 

The accuracy of the measuring instruments and test equipment shall be verified periodically by calibration. 

The maximum environmental tolerances (except instrumentation tolerances) on test conditions during 

environmental testing shall be as specified in Table 6.1-2 unless otherwise specified. 




Parameter Measurement Range Tolerances 

Mass + 0.010 kg or 0.15% whichever is greater 

Volume + 0.06 dm3 Temperature maximum temperature 

+ 3°C 

minimum temperature 

- 3°C 

Pressures system pressure p > 1bar 

+ 1 % full scale 

p < 1 bar 

+ 2 % full scale 

barometric pressure: p > 0.1 mbar 

+ 5 % 

p < 0.1 mbar 

+ 50 % 

Measured pressure above 60% of full scale. 

Relative Humidity + 3 % RH 

Acceleration + 10 % 

Vibration sinusoidal 

+ 10 % g-peak 

random PSD 20 - 300 Hz 

+ 1.5 dB 

300 - 2000 Hz 

+ 3 dB 

random rms 

+ 10 % 

Frequencies < 20 Hz 

+ 0.5 Hz 

> 20 Hz 

+ 5 % 

Time + 1 % 

Sweep Rate + 5 % 

Acoustic Pressure ± 3dB per octave band, ±1.5dB OASPL 

Force static tests + 5% / - 0 % 

Force and Moments dynamic tests + 10 % 

Leakage Rate + 50 % 

Mass Flow + 10 % 

Graduated Cylinder + 1 % 

Electrical Conditioning Voltage < 5 Volt 

+ 0.2 % 

> 5 Volt 

+ 0.5 % 

Current + 0.1 % 

Resistance high 

+ 10 % 

low 

+ 2 % 

Centre of Mass Deviation from nominal centre + 0.5 mm 

Moments of Inertia Measurements, if MOI > 0.1 kgm2 

+ 10% 

Calculations, if MOI < 0.1 kgm2 

+ 10% 

Table 6.1-2 : Measurement requirements 




6.1.4.3 Cleanliness of test equipment 


PL - 6.1.4 -4 

The inner cleanliness of the test equipment as far as it can affect the cleanliness of the Payload shall be 

checked and minimum cleanliness level shall be assured before, during and after each test. 

6.1.4.4 Measurements 

PL - 6.1.4 -5 

During all tests to be performed, the test data and parameter values shall be continuously recorded. 

PL - 6.1.4 -6 

Prior to conducting any of the tests, the test item shall be operated under ambient conditions, and a record 

shall be made of all data necessary to determine compliance with the required performance in the 

subsequent performance tests conducted before, during and after the environmental exposure. The only 

exceptions to this requirement are for those items which cannot be tested realistically in ambient conditions. 

In such cases, initial testing shall be designed to prove compliance as far as possible without causing 

damage to the test item. 




6.1.5 Electrical Verification 

6.1.5.1 Electrical Functional Verification 

PL - 6.1.5 -4 

Three kinds of payload reference tests shall be defined and conducted during payload functional tests and 

before payload delivery 

These tests will serve as a baseline against which all later results can be compared. The results obtained 

during the satellite tests shall be similar to the ones obtained during payload acceptance. 

These tests are: 

• Payload Performance Verification Test (PPVT) 

• Payload Health Check Test (PHCT 

• Payload Aliveness Test (PAT) 

PL - 6.1.5 -5 

The duration of these tests defined in following sections shall be optimised to the strict necessary. All the 

operations shall be individually justified. 

6.1.5.1.1 Payload Performance Verification Test (PPVT) 

PL - 6.1.5 -1 

A reference Payload Performance Verification Test shall be conducted at payload level before payload 

delivery. This reference PPVT shall be a part of the total PPVT and will serve as a baseline against which all 

later results can be compared. The results obtained during this part of PPVT shall be similar to the ones 

obtained during Payload acceptance. The duration of this part of the PPVT shall be lower than 5 days (i.e. 5 

x 8 hours). 

The total PPVT shall be a detailed demonstration that the hardware and software meet all their performance 

requirements within allowable tolerances. It shall exercise all Payload modes, science operations and 

calibration measurements (where applicable). It shall also demonstrate operation of all prime and redundant 

components and hardware (where applicable) and shall be performed for each Payload side (where 

applicable). 

6.1.5.1.2 Payload Health Check Test (PHCT) 

The Payload Health Check Test is a subset of the PPVT. 

PL - 6.1.5 -2 

The Payload HCT shall exercise major Payload modes, limited science operations and calibration 

measurements (where applicable). 

The Payload HCT shall demonstrate for each Payload side operation of all prime and redundant 

components and hardware (where applicable). 

The HCT is a part of the PPVT. 

6.1.5.1.3 Payload Aliveness Test (PAT) 

PL - 6.1.5 -3 

The Payload Aliveness Test shall test engineering housekeeping functions only; no Payload science testing 

will be performed. 

The Payload AT shall be performed during short duration to check the communication between the DHU and 

the Payload. 

The duration of PAT shall be lower than 1 day (i.e. 8 hours). 




6.1.5.2 Electrical Interfaces Validation 


The Payload Supplier is in charge of demonstrating the payload compliance with regard to all I/F requirements by 

analyses, similarity or tests. When tests are proposed, the payload responsible shall detail the configuration, the way 

how the test is intended to be performed and the success criteria. 

In order to somehow clarify the satellite contractor needs, the following sections give some requirements for this 

validation. 

6.1.5.2.1 Platform interfaces simulation 

PL - 6.1.5 -6 

If an EGSE is specifically developed to simulate electrical I/F, it shall be fully representative (in term of 

electrical characteristics and grounding) of the platform interfaces as described in the section 3.5. 

Moreover, implementation in the EGSE of PF I/F electrical schematics & layouts as defined in Appendix E is strongly 

recommended. 

PL - 6.1.5 -7 

The length and the type of harness used between EGSE and payload shall be representative of the platform 

harness. Demonstration of the EGSE representativeness shall be provided by the payload responsible 

(compliance matrix with regard to platform characteristics shall be provided. 

PL - 6.1.5 -8 

The electrical validation shall be performed in configuration representative of the flight hardware 

configuration. In particular, the EGSE/payload overall grounding network shall be fully representative of the 

satellite grounding. Demonstration of the representativeness of the grounding network shall be provided by 

the payload responsible. 

6.1.5.2.2 Electrical interface validation 

PL - 6.1.5 -9 

Tests shall be done with the payload fully integrated. The measurements shall be done at connector/bracket 

levels. 

PL - 6.1.5 -10 

All signals shall be tested. 

6.1.5.2.2.1 Pin allocation/signal addressing 

PL - 6.1.5 -11 

All signals shall be addressed during payload test. 




6.1.5.2.2.2 Continuity/Isolation 

PL - 6.1.5 -12 

The following isolation or continuity shall be checked: 

• Isolation between each primary power line & P/L structure. 

• Isolation between each heater line & P/L structure. 

• Isolation between each pyro line & P/L structure (real pyro initiator replaced by dummy). 

• Isolation between each DR line & P/L structure. 

• Isolation between each TH line & P/L structure. 

• Isolation between each HLC line & P/L structure 

• Isolation between each LLC line & P/L structure 

• Isolation between each 1553 line & P/L structure (long stub case). 

• Continuity between each signal line connected to secondary zero-volt line (inside a unit) e.g. ANA/DB 

return and P/L structure. 

6.1.5.2.2.3 Primary Power lines 

PL - 6.1.5 -13 

The following measurements shall be performed: 

• In-rush current measurement (covered by payload EMC test) 

• permanent current measurement for each functional mode (P/L power consumption budget 

consolidation), 

• other tests for PF to PL electrical interfaces validation are covered by payload EMC test, isolation and pin 

out verification. 

6.1.5.2.2.4 Heater lines 

Heater lines interfaces validation are covered by payload EMC test, isolation and pin out verification. 

6.1.5.2.2.5 Other lines 

PL - 6.1.5 -14 

For other electrical lines, the requirements for validation are provided in the following table. 




Signal type Voltage Timing Waveform Impedance Triggering 

Comments 

level 

(Tr, Tf, pulse 

Threshold 

duration, …) 

/Hysteresis 

DM CM 

ANA X X * 

DB X X * 

DR X ** Open circuit & Close circuit to be checked 

Th X ** Measurement at ambient temperature 

HLC X ** X Functionality to be validated over the EGSE 

active voltage level range and over Tr and Tf 

pulse range 

LLC X ** X Functionality to be validated over the EGSE 

active voltage level range and over Tr and Tf 

pulse range 

PPS X 

ML16 C/E/D X X X X X ML16 & DS16 differential receivers 

& DS16 C/E 

DS16 D X X X X X DS16 data transmitter 

1553 X X X X ** 

X : to be measured * : covered by continuity test ** : covered by isolation test 

Table 6.1-2b: Payload electrical interface verification matrix 

Nota : the compliance the fault tolerance requirements of Payload interface shall be demonstrated by analysis for 

instance. 




6.1.6 Structural and Mechanical Verification 

6.1.6.1 Sinusoidal Vibrations 

The sinusoidal vibrations test aims at demonstrating the capability of the test item to withstand and properly function 

after the sine vibration environment encountered during launch and of the test item primary structure to withstand the 

quasi-static loads encountered during launch. 

This test may also reveal defects in design, parts, and workmanship, if any. 

And finally, this test allows to demonstrate that the structural design of the test item shows the proper response to 

sine excitation. In particular, it identifies the critical lowest resonance frequency of the item. 

PL - 6.1.6 -1 

The Payload shall undergo a Sinusoidal Vibrations Test before payload delivery. 

PL - 6.1.6 -2 

The levels given in Table 5.1-2 shall be used for sine vibration design qualification (levels are 0-to-peak). 

Acceptance levels are 1.25 times lower (launch vehicle dependent). 

Qualification testing shall be with a sweep rate of 2 octaves per minute, one sweep up. 

Acceptance testing shall be with a sweep rate of 4 octaves per minute, one sweep up. 

For Protoflight testing, see PFM definition section 6.1.2.3. 

Notching philosophy at payload level is defined in section 4.2.5.3. 

For information, notching at any instrument vibration frequency will not be allowed. 

PL - 6.1.6 -3 a 

Before and after sine test along each axis at the qualification levels required in Section 5.1.2, a low level sine 

test (0.5 g from 5 to 2000 Hz, 2 octaves/min, 1 sweep up) shall be performed with the objective to 

demonstrate that the unit has not been damaged by the qualification test. 

The sine vibrations tests are usually performed on a shaker along all three axes in sequence. 




6.1.6.2 Acoustic/Random Vibrations 


These tests aim at demonstrating its ability to survive mechanical stresses arising during pre-launch and launch 

environments. 

They will also reveal defects in design, parts, and workmanship, if any. 

As a general rule, acoustic test applies to big size payload whereas random vibrations test applies to smaller one. 

PL - 6.1.6 -4 

The Payload shall undergo an Acoustic or Random Vibrations Test before payload delivery. 

PL - 6.1.6 -5 

The levels given in Table 5.1-3 or Table 5.1-4 shall be used respectively for random and acoustic vibrations 

design qualification. 

Qualification testing shall be through a test duration of 120 s on each axis. 

Acceptance testing shall be through a test duration of 60 s on each axis. 

Test Tolerances for the sound pressure levels are given Table 6.1-3. 

Octave Band Center Frequency 

(Hz) 

Test Tolerances 

(dB) 

31.5 -2/+2 

63 -1/+2 

125 -1/+2 

250 -1/+2 

500 -1/+2 

1000 -1/+2 

2000 -1/+2 

4000 -3/+3 

8000 -4/+4 

Overall -1/+3 

Table 6.1-3 : Acoustic test tolerances 

The random vibrations tests are usually performed on a shaker along all three axes in sequence. 




6.1.6.3 Shocks 


Shock tests aim at demonstrating the capability of the test item to withstand and properly function after the shock 

environment encountered during and after launch (satellite separation, solar arrays deployment). 

They will also reveal defects in design, parts, and workmanship, if any. 

As a general rule, shock testing is not required for structural components. 

PL - 6.1.6 -6 

The Payload shall undergo a Shock Test at payload or payload sub-system level (payload dependent) before 

payload delivery. 

The shock response spectrum is specified in section 5.1.5. 

PL - 6.1.6 -7 

The Payload shall verify by test that it does not generate shock levels higher than those given in section 

3.1.5.2. 

6.1.6.4 Handling 

PL - 6.1.6 -8 a 

Tests at payload level shall be performed in order to demonstrate the possibility of handling the payload at 

ALCATEL SPACE facilities. These tests shall be performed before delivery and with additional masses in order 

to represent the maximal payload masses (see § 4.2.2.4). 

The maximum load encountered during nominal handling shall be tested on the flight hardware. 

The maximum load encountered during degraded case (fail-safe for instance) shall be tested on a 

representative sample. 

For safety reasons, related test reports shall be provided with the payload for its acceptance. 




6.1.7 Thermal Verification 

PL - 6.1.7 -1 

The thermal active control of the payload (including the lines provided by the platform) shall be qualified at 

payload level before delivery to the satellite (thermal balance test). 

The regulation parameters (C1, C2 and Tref) shall also be adjusted before payload delivery. 

That involves that only minor modifications of the regulation parameters will be authorised during satellite thermal 

tests. 

6.1.7.1 Thermal Balance Test definition 

A test conducted to verify the adequacy of the thermal model, the adequacy of the thermal design, and the capability 

of the thermal control system to maintain thermal conditions within established mission limits. 

6.1.7.2 Thermal Vacuum Test (Thermal Cycling) definition 

A test to demonstrate the capability of the test item to operate satisfactorily in vacuum at extreme temperatures based 

on those expected for the mission with adequate margin. The test can also uncover latent defects in design, parts, 

and workmanship. 




6.1.8 EMC verification 

The test sequence applies to the unit level tests, which are to be analysed for earliest possible prediction of whatever 

problems may arise. Prediction of the units test results, prior to the tests, shall form the input for writing the Test Plan, 

with a view to prevent any potential incompatibility between the units and evidence shortcomings, if any, in the 

compatibility margins. 

These tests aim at demonstrating the capability of the test item to operate satisfactorily under the electromagnetic 

environment encountered during the mission. The test also demonstrates that the test item does not generate more 

electromagnetic interference than specified. 

6.1.8.1 Test Configuration 

The tested hardware, whether at overall satellite or at Payload level, shall be confronted with all operational modes 

for which it was originally designed, with each synchronizable converter operating in synchronized mode. 

PL - 6.1.8 -1 

The complete EMC tests could be run on one single model of the Payload if several replicas are built. This 

test article shall be fully representative of the Flight Model (if not, the full test sequence shall be run on the 

FM). Furthermore, where a change has occurred from Engineering Model (EM) or Qualification Model (QM) 

to Flight Model (FM), the EMC tests shall be performed again following the same approach. 

PL - 6.1.8 -2 

A reduced EMC tests sequence will be done with the FM in case the complete EMC tests have been 

performed on another model. 

PL - 6.1.8 -3 

In all operating modes, the aim of the EMC tests shall be a worst-case assessment of both emission and 

susceptibility. 




6.1.8.2 Test Requirements 


6.1.8.2.1 Units and harness/wiring configuration 

PL - 6.1.8 -4 

The units and harness configuration shall be as 

• the lay-out of units is the nominal lay-out, 

• actual (flight-standard) inter-unit or inter-connector harness/wiring at final positions, 

• interface connectors linked to simulators or to dummy loads. 

6.1.8.2.2 Test operating conditions 

PL - 6.1.8 -5 

Test harness/wiring and Payload units shall not create any grounding loops. 

PL - 6.1.8 -6 

If some radiated perturbations are measured in an anechoic test room, the simulators and EGSEs shall be 

located outside, and the level of ambient perturbations shall be at least 6 dB below the required level. 

PL - 6.1.8 -7 

The test-operation and parameter-measuring procedures shall be that applicable to the system to be tested. 

PL - 6.1.8 -8 

Should, in any test sequence, an operating mode appear as the most unfavorable in terms of EMC, that 

mode shall be selected. 




6.1.8.2.3 Band analyses 

PL - 6.1.8 -9 

The analysis bandwidth to be used for Narrow band (NB) and Broadband (BB) recordings shall be as 

follows: 

TYPE FREQUENCY RANGE BANDWIDTH 

NB up to 10 kHz < 50 Hz 

NB 10 kHz - 2.5 MHz < 500 Hz 

NB 2.5 MHz - 25 MHz < 5 kHz 

NB 25 MHz - 1GHz < 50 kHz 

NB 1GHz - 10 GHz


6.1.8.4 Test Site 


PL - 6.1.8 -28 

The Payload shall be placed within an anechoic chamber whose walls are coated with absorbing panels, to 

achieve a reflecting coefficient less than -20 dB in the 100 MHz to 18 GHz frequency band. Use of such 

anechoic chamber may not be required if the following conditions are demonstrated for the selected test site: 

• - 20 dB reflection coefficient achieved, at the Payload transmission frequencies, by other means, 

• ambient noise level less than 6 dB below the levels defined herein; ambient noise may locally exceed the 

limits set if occurring as predictable, stable, discrete frequencies sufficiently spaced throughout the 

frequency band. 

6.1.8.4.1 Facility requirements 

PL - 6.1.8 -11 

Unless from out-of-control impossibility, the tests shall be run within an anechoic chamber, which shall 

comply with the following prescriptions: 

• dimensions shall be such that antennas always stand at a distance not less than 1 m to any of the test 

chamber walls, except for the dipole and whip antenna, for which the minimal distance can be reduced 

to 30 cm. 

• filtering of the power sources shall curb residual perturbations to less than the limits set hereby by at 

least 6 dB, with the ambient electric fields lower by at least 6 dB than the limits set hereby; for the 

purpose of those measurements, the power source shall be closed on a charge at least equal to the 

tested unit, with the measuring Payloads and test devices ON. 

• However, should the global level, i.e. ambient perturbations added to the perturbations on the tested 

unit, lie within the limits set, the equipment shall be ruled satisfactorily. 

• The chamber shall feature a grounding plane as per MIL.STD.462, at least 2 m 2 in size and 0.75 m in 

width, which may be formed by a copper, brass, or lightweight alloy plate of a minimal thickness of 0.70 

mm; the plate connections to the chamber, preferably made up of 0.7 mm thick copper strips whose 

width equals at least 20% of their length, shall be spaced by no more than 1 m. 

Use of absorbing materials with anechoic properties is recommended. 

6.1.8.4.2 Tests outside an anechoic chamber 

The selected test site shall be a clear area, free from rough/uneven features, preferably located near an 

earthling/grounding outlet linking up to the grounding plane, which may consist of a copper, brass or lightweight 

alloy not less than 2 m 2 in size, with ambient EMC perturbation control identical to that defined here before. 




6.1.8.4.3 Measuring instrument 

6.1.8.4.3.1 Measuring receiver 

The minimal characteristics required of measuring receivers are as listed below: 

input impedance: 50 Ω 

input TOS: 

< 1.5 up to 200 MHz 

< 2 from 200 MHz to 18 GHz 

recommended analysis bands: 

see 6.1.8.2.3 (If needed, the analysis bands shall be reduced to decrease measurement noise). 

types of detection recommended: 

. peak, efficient, mean 

accuracy of frequency measurement: 1% 

accuracy of voltage: 2 dB 

6.1.8.4.3.2 Spectrum analyser 

The spectrum analyser may be used as a measurement receiver, using analysis bands similar to those specified for 

the receiver. 

6.1.8.4.3.3 Electric field measurement antennas 

Depending on measuring frequencies, the following antennas may be used: 

- below 30 MHz: whip antenna, 1 m in length. 

- from 20 MHz to 200 MHz: dipole antenna 

biconical antenna 

- from 200 MHz to 1 GHz: dipole antenna 

‘log spiral’ conical antenna 

so-called ‘ridged guide’ antenna 

- from 1 GHz to 12.4 GHz: ‘log spiral’ conical antenna 

so-called ‘ridged guide’ antenna 

- beyond 10 GHz: horn- shaped, parabolic-shaped antennas. 

This list is not restrictive: any other antenna may be used, provided its "antenna factor" at receive/transmit stage is 

known from calibration or from the manufacturer’s diagrams. 

6.1.8.4.3.4 Calibration 

PL - 6.1.8 -29 

The last-calibration date for the measuring instruments used in the EMC tests shall be less than one year old. 

This requirement does not apply to the passive (current probe-type) instruments, as those shall exhibit a 

calibration curve. 




6.1.8.4.4 Test set-ups 

PL - 6.1.8 -12 

Generally speaking, the test conditions and set-ups shall comply with standard MIL.STD.462 where 

applicable. 

PL - 6.1.8 -13 

Moreover, the Payload to be tested shall be installed in conditions that best reproduce the normal conditions 

of use, particularly at unit level: 

• grounding, shielding and backshell identical to conditions of use 

• harness/wiring of same nature and immunity as at unit installation 

• accurate definition of Payload-associated harness/wiring shall be provided 

• the antenna of every tested receiver or transmitter shall be replaced with a dummy antenna or with a 

shielded charge of equivalent impedance to that of the actual antenna. 

PL - 6.1.8 -14 

Both test conditions and test set-ups shall be described in the Test Plan. 




6.1.8.5 TESTS 


6.1.8.5.1 Conducted test requirements 

PL - 6.1.8 -15 

The requirements given in section 3.5.7 shall be verified at a power supply voltage of 37 V for conducted 

emission and 23 V for conducted susceptibility. 

PL - 6.1.8 -16 

A Line Impedance Stabilised Network (LISN, Figure 6.1-1) shall be used to simulate impedance of primary 

power supply. The wound, unshielded test connections shall be with the negative power supply point on LISN 

input grounded. 

PL - 6.1.8 -17 

The LISN shall be used in each EMC test except if mentioned . Its characteristics shall be measured by the 

Instrument Contractor and delivered with the EMC test report. 

PL - 6.1.8 -18 

The star point of the test set up shall be in the LISN, on the return link. 

R1, R2 < 20 mOhm (inductance parasistic resistors) 

R3, R4 = 50 Ohms 

L1, L2 = 4 µH 

C1 = 19 mF 

Figure 6.1-1: Schematic representation of a LISN 

PL - 6.1.8 -19 

As far as possible, Payload shall be tested within a shielded chamber. The test requirements shall be as 

follows: 

• Conducted emission and susceptibility to be tested on Engineering or Qualification model (EM or QM) 

equipment. 

• Need for partial or full tests at Flight Model (FM) level to be analysed in case of changes and where 

components technology or manufacturer are not identical for EM or QM and FM. 




Flight Model (FM) tests consist in: 


measuring inrush current at time-domain on switching, 

measuring conducted emissions in both common and differential modes. 

All the above involve both the nominal and the redundant channels. 

Identification Measuring range Section N° 

Conducted emissions (CE) 10 Hz - 50 MHz 

Power supply bus 

Wave 3.5.7.1.1 

Transients 

Emissions from transmitter units/devices 10 kHz - 18 GHz 

Conducted susceptibility (CS) 10 Hz - 50 MHz 

Power supply bus 

Sine wave 

Modulated wave 3.5.7.1.2 

Transients 

Receiver units/devices 10 kHz - 18 GHz 

Table 6.1-5: Measuring range and section for CE and CS tests 




6.1.8.5.2 Radiated test requirements 

Identification measuring range Section N° 

Radiated emissions (RE) 

Non-RF equipment 10 kHz - 1 GHz 3.5.7.2.1 

RF equipment 10 kHz - 18 GHz 

Radiated susceptibilities (RS) 

Sine 10 kHz - 1 GHz 3.5.7.2.2 

Modulation 10 kHz - 18 GHz 

Table 6.1-6: Measuring range and section for RE and RS tests 

6.1.8.5.3 Electrical Ground Support Equipment (EGSE) 

Regarding EGSE, the requirements here before are applicable only to those EGSE to be co-located with the Payload 

during radiated emission/susceptibility tests. 

PL - 6.1.8 -20 

Such EGSE, complemented with the dedicated harness/wiring interfacing the tested unit, shall be subjected to 

the following measurements: 

• Narrowband radiated emissions over the 10 kHz to 18 GHz range (up to 1Ghz for all units except RF 

units), 

• Susceptibility to the Payload transmit frequencies. 




6.1.8.6 Tests organization 


6.1.8.6.1 Test Plan 

PL - 6.1.8 -21 

Each electromagnetic compatibility test shall be defined in a dedicated document drawn up by the Payload 

Supplier. This document, which constitutes the Test Plan, contains the specific data to be used in writing the 

test procedure. 

The following pieces of information shall be provided in the Test Plan: 

a) Specimen operational configuration during test, 

b) Duly justified choice of a measuring method, 

c) The bare descriptive modicum for environmental and operational conditions, 

d) Specimen operating modes and points to be watched (susceptibility criteria), 

e) Description of injected signals for measuring susceptibility or the compatibility margin. 

6.1.8.6.2 Test procedure 

The unit development, qualification or verification EMC tests follow a test procedure that details how tests must be 

run to verify compliance with the EMC requirements. 

PL - 6.1.8 -22 

The procedure shall be made available to the Satellite Contractor for approval one month at least before test 

inception, and shall contain at least the following sections, in sequential order: 

a) contents, 

b) applicable documents, 

c) purpose of tests, 

d) general test conditions (electromagnetic environment, grounding plan, measuring precautions, authorized 

personnel, power supply characteristics), 

e) specimen detailed mechanical and electrical configuration (operating mode, power supply voltage, input 

signals, stimuli, dummy charge power levels, points to be watched, detailed description of interface 

harness/wiring, overall layout on test site, grounding connection), 

f) for each type of test: 

• required test instrumentation, 

• antenna calibration data sheets, 

• measurement set-up, accuracy over the specific precautions for each type of test, 

• test limits and levels, 

• frequency ranges or discrete frequencies for the test, 

• susceptibility criterion. 




6.1.8.6.3 Test execution 

PL - 6.1.8 -23 

Test execution shall be documented by the proceedings from the test sequence as actually experienced, 

which state the facts as observed in real time: 

a) calibration of the actually used instruments, 

b) recording of measurements (photos, plots, graphs, tables, etc.), 

c) deviations from procedures, or changes required by real conditions. 

PL - 6.1.8 -24 

Assessment of the specimen compliance with the specifications shall be acquired during tests, with a clear 

identification of non-conformance, e.g.: 

a) measurement of emitted level, should the emission limit be exceeded 

b) measurement of susceptibility threshold, if actual threshold is less than specified, 

c) measurement of real compatibility margin, if found less than specified. 

The above elements are critical to obtaining a waiver for not meeting a specified requirement. 

6.1.8.6.4 Presentation of results 

6.1.8.6.4.1 General 

The rough results shall be of the following form: 

XY recording with sufficient resolution to ease out data analysis, 

photos of oscilloscope or spectrum analyser. 

The data recorded in emission tests shall be read in continuous frequency-scanning mode. 

For a quick assessment of results, the test data and the maximum levels allowed by the present specification 

(requirements of section 3.5.7) shall be presented on the same plots. 

All such auxiliary data as sensitivity, bandwidth, antenna factor, aso., shall be provided along with the data and 

photos. 




6.1.8.6.4.2 Test report 

PL - 6.1.8 -25 

The unit EMC test report shall be submitted to the Satellite Contractor by the Payload Supplier within thirty 

(30) days from official completion of the EMC tests, complete with the relevant test procedures. 

PL - 6.1.8 -26 

For uniformity, and to ease out analysis, such test report shall contain at least the following: 

a) Contents 

b) Purpose of test 

c) Changes to nominal procedure 

d) Summarized results 

e) Conclusions 

f) Working copy of the procedures, containing: 

• description of test set-up, with photos of test configuration, 

• detailed description of grounding network, 

• problems encountered and corrective actions, 

• type and serial number of the measuring instrumentation, date of last calibration, 

• measures of ambient noise including EGSE, 

• raw measurement sheets, recordings, 

• transfer function of actually used probes or antennas, 

• interpretation of measurements against specified noise, 

• complementary measurements performed, as applicable. 

PL - 6.1.8 -27 

In addition, the test report shall spell out and justify all deviations from, or changes to the test procedure, 

which procedure shall have been approved by the Satellite Contractor prior to official inception of the EMC 

tests. 




6.1.8.7 Unit Test set-ups 


6.1.8.7.1 Conducted emissions; Power supply lines, steady perturbations 

Methods: CE01 - CE04 of MIL-STD-462 

Test set-up: 

1. 5 cm Stand-off 

2. Low-impedance bond to ground plane 

3. Current probe 

4. Test sample chassis ground 

5. High side 

6. Return (neutral line) 

7. DC bond impedance between the ground plane and enclosure wall 

8. Line impedance stabilization network shall be terminated in 50 Ohm resistive. 

Figure 6.1-3 : Power lines, steady perturbations test set up 




6.1.8.7.2 Conducted emissions; Power supply lines, transient perturbations 

Method: out of standards 

Test set-up: 

Test set-up is identical to that one used in steady perturbation measurements, except that an oscilloscope is 

substituted for the spectrum analyser, and that the current probe bandwidth has to be adapted to the measurement 

signal. 

Figure 6.1-4 : Power supply line, transient perturbations test set-up 

Observations : Whatever measurements are needed over the command lines at the switching unit, outputs are made 

with representative harness/wiring and charges, without any LISN or 10 µF capacities. 




6.1.8.7.3 Conducted susceptibility ; power supply lines, sine wave and square wave 

Methods: CS01 and CS02 of MIL-STD-462 

Test set-up: 

CS01, Differential mode 

Generator 

Measur. Inst. 

Current probe 

Power supply = LISN V Tested unit/device 

CS02, Common mode 

Measur. Inst. 

Power supply = LISN Tested unit/device 

10µF 

Generator ~ V V 

Figure 6.1-5 : Conducted susceptibility test set-up (sine wave and square wave) 




6.1.8.7.4 Conducted Susceptibility; power supply lines, transient signal 

Method: CS06 of MIL-STD-462 

Test set-up: 

Power supply = LISN Zo Tested unit/device 

CS, spike, power leads, injection in series 

Pulse generator Oscilloscope 

L = 20µH 

Filter 

Power supply= LISN Tested unit/device 

Pulse 

Generator Zo 

CS, spike, power leads, injection in parallel 

Filter Oscilloscope 

Figure 6.1-6 : Conducted susceptibility test set-up (transient signal) 




6.1.8.7.5 Susceptibility to common mode transients; interface signals 

Method: out of standards 

Test set-up: 

EGSE 

Generator 

Intermediate connector 

disconnecting shields 

Figure 6.1-7 : Interface signals test set-up 

Unit/device under 

test 




6.1.8.7.6 Radiated emissions E-fields 

Test method: RE02 of MIL-STD-462 

Test set-up : 

Figure 6.1-8 : Radiated emissions E-fields test set-up 

The unit/device to be tested is installed and connected to the ground plane. 

Harness/wiring of the tested equipment shall be flight representative: same type, same twisting, same gauge, same 

shielding connection mode as on the flight model. 

Such harness/wiring shall be kept 2 to 3 cm clear above the ground plane, and shall as far as possible offer a length 

of 1 meter maximum, 1 meter off the measurement antenna. 

Remarks : 

Measurements shall be made in two linear cross-polarizations, or in one circular polarization beyond 50 MHz. 




6.1.8.7.7 Radiated susceptibilities E-fields 

Method :RS03 of MIL-STD-462 

Test set-up: 

Figure 6.1-9 : Radiated susceptibilities E-fields test set-up 

The unit to be tested is installed and attached to the ground plane. 

All of the tested equipment harness/wiring (not the power supply strands only) shall be representative in their nature 

of the real, flight-standard harness/wiring: same type, same twisting, same gauge, same shielding connection mode. 

Lengths shall be limited to 2 meters for significant lengths. 

Such harness/wiring shall be kept 2 to 3 cm clear above the ground plane, and shall as far as possible offer a length 

of 1 meter, 1 meter off the measurement antenna. 

Remarks : 

Such test sample orientation shall be sought to maximize emitted perturbations. 

Measurements shall be made in two linear cross-polarizations (or in one circular polarization beyond 50 MHz). 

6.1.8.7.8 Magnetic moment (DC) 

The measuring procedure to be used shall be subject to prior approval by the Satellite Contractor. 




6.1.9 ESD Verification 

PL - 6.1.9 -1 

The payload shall verify its compatibility with the arc discharge described in section 3.5.8. 

PL - 6.1.9 -2 

The test shall be performed with the discharge electrodes being directly applied on the payload chassis and 

cables shields for repetitive electrostatic discharges of 10 to 15 kV. 

PL - 6.1.9 -3 

The repetition rate shall be 1 ESD pulse per second, during at least 3 minutes. 

6.1.10 Magnetic field Verification 

PL - 6.1.10 -1 

If there are magnetic elements, Magnetic Cleanliness control shall be performed at unit level. It shall include 

the following: 

• quality control of parts and material including magnetic characterization, 

• magnetic characterization test in a dedicated magnetic facility. 




6.1.11 Verifications prior to Payload Delivery 

6.1.11.1 Inspections and examinations at unit level 

PL - 6.1.11 -1 

The Payload shall be examined to verify compliance with the following criteria: 

• Configuration, 

• Interface Requirements, 

• Parts, Materials and Process, 

• Identification and Marking, 

• Workmanship. 

6.1.11.2 Mass properties determination 

PL - 6.1.11 -2 

The mass shall be determined by weighing before delivery for satellite integration. 

PL - 6.1.11 -3 

The moments of inertia and the center of gravity shall be determined by test. 

6.1.11.3 Unit acceptance and delivery for satellite integration 

PL - 6.1.11 -4 

After completion of all acceptance tests at Payload level, accepted Payload shall be appropriately sealed by 

the Payload Supplier QA and released for storage or transportation or integration. 

An Acceptance Data Package as defined in Payload Deliverable Items List shall be delivered with the unit. 




6.2 TESTS AND VERIFICATIONS AT SATELLITE LEVEL 

Figure 6.2-1 shows the main sequence of assembly, integration and test at satellite level. 

Figure 6.2-1 : Satellite Assembly Integration and Test 




6.2.1 Payload Inspection before Integration 

PL - 6.2.1 -1 

The Payload shall be examined visually to verify that no handling damage has occurred. 

6.2.2 Functional tests 

The Functional tests will be as follows : 

What kind ? How many ? When ? 

AT 3 Initial, final & post vibration 

HCT 2 EMC & thermal vacuum 

PVT 2 Initial & final 

Table 6.2-1: Functional tests 




6.2.3 Thermal Vacuum tests 

The satellite will be submitted to a thermal balance test (cold case) to correlate the platform thermal mathematical 

models and qualify satellite thermal control and to a thermal vacuum test (thermal cycling) to verify the satellite 

ability to meet the qualification requirements under vacuum conditions and extreme temperatures, which simulate 

those predicted in flight with qualification margins. 

To obtain steady case temperature a vacuum chamber with liquid nitrogen-cooled shroud is used. The pressure will 

be lower than 10-4 torr. The thermal conditions will be established with infrared heat sources or by skin heaters. The 

choice could be different for the platform and for the payload. Nonetheless, with regard to the payload, this thermal 

environment shall be simulated thanks to the dedicated electrical facilities specified in PL-6.2.3-4. 

Moreover the Payload shall be compatible with the « ESPACE 70 » thermal vacuum facility of Alcatel Space Cannes. 

Consequently, the payload shall comply with the following requirements. 

PL - 6.2.3 -1 

The payload and its specific thermal facilities and instrumentation in thermal vacuum test configuration must 

be enclosed in a volume defined as a cylinder centred on the Satellites Xs axis with a diameter less than 3.60 

m and a height less than 2.80 m. 

PL - 6.2.3 -2 

The mechanical configuration of the payload during thermal vacuum test shall be, indiscriminately, the 

following : 

• even Xs horizontal and Ys vertical (Satellite axis) 

• even Xs horizontal and Zs vertical (Satellite axis) 

PL - 6.2.3 -3 

The Payload shall have its own thermal thermocouple instrumentation. The maximum allocation for Payload 

thermocouples is : 

• 110 thermocouples in the Cu/Cs class 

• 25 thermocouples in the Cr/Al class 

PL - 6.2.3 -4 

If needed, the Payload shall have its own thermal facilities dedicated to external fluxes simulation in orbital 

environment (solar, albedo and IR earth fluxes). In order to achieve this simulation, the Payload has at its 

disposal, during test, the following maximum electrical power lines allocation : 

• 3 lines in the category : { Pmax=100W under Umax=50V with Imax=2A } 

• 2 lines in the category : { Pmax=200W under Umax=35V with Imax=5.5A } 



These electrical power lines are driven by a dedicated computer. 

Note that the programming of a succession of different required power for each line during the test is 

possible (by modifying the needed power at each step of update of the command, for instance each 30 or 

60 seconds). 




PL - 6.2.3 -5 


The satellite thermal test campaign is divided in 2 phases : 

• first, a thermal balance test in order to declare the final qualification of the satellite 

• second, a functional test in simulated spatial thermal conditions in order to qualify and verify 

performances of the satellite in extreme thermal environment configurations. 

These tests are performed in succession, without chamber pressure or shroud temperature modifications. 

The maximum duration for satellite thermal test campaign is : 

• 3 days for the thermal balance test step 

• 10 days for the functional verification in thermal environment condition step. 

PL - 6.2.3 -6 

During all the tests, thanks to the efficient thermal uncoupling between the payload and platform, no 

constraint on thermal configuration synchronisation is required between the payload and the rest of the 

satellite. Nevertheless, the PL thermal environment simulation facilities (in case of use of infrared heat 

sources) shall not generate fluxes towards the platform. 

PL - 6.2.3 -7 

During the thermal balance step, the functioning scenario of the payload units (with the exact timing) is 

defined by the payload and specified to the satellite (if any). 

During the functional test in simulated spatial thermal conditions step, the functioning scenario of the 

payload units is determined in accordance with the satellite. 

This information shall be supplied at the latest 4 months before the beginning of the satellite thermal test 

campaign. 

PL - 6.2.3 -8 

The Payload shall supply ALCATEL SPACE with the detailed mechanical, thermal and electrical ICD (Interface 

Control Documents) and IDS (Interface Data Sheets) of the payload in its thermal satellite test configuration, 

including instrumentation and dedicated thermal test facilities, at the latest 6 months before the beginning 

of the satellite thermal test campaign. 

PL - 6.2.3 -9 

The Payload shall supply ALCATEL SPACE with all the data allowing the monitoring of the thermal test. More 

particularly, the Payload must deliver all the parameters of the thermal test facilities dedicated to the payload 

and monitored by ALCATEL SPACE during all the tests (thermal regulation parameters, test heaters 

instructions, ...). 

This information shall be supplied at the latest 2 months before the beginning of the satellite thermal test 

campaign. 

The vacuum before first turn on shall be 10-5 hPa. During cycling, temperature and current will be continuously 

monitored. During these tests, provisions will be made to prevent the Payload from exceeding the specified operating 

temperature limits. 




6.2.4 Vibration tests 

PL - 6.2.4 -1 

The Payload shall have its own mechanical instrumentation. The maximum allocation for Payload 

instrumentation is: 

• 50 sensors for sine, acoustic (or random) vibrations and shock tests (no specific instrumentation is 

foreseen for shock tests). 

6.2.4.1 Sinusoidal Vibrations 

The satellite will be qualified for each of the three axes with the following sequence: 

Low level sine scan for resonance search (verification of the primary notching if any). 

Intermediate level run, with notching (notching defined divided by 2), performed with qualification levels 

divided by 2. 

Qualification level, with notching defined, at acceptance sweep rate. 

Control low level (to verify that vibration did not modify the behaviour of the satellite). 

To avoid unrealistic overtesting, the sine spectrum may be adjusted by notching the input on the basis of the load 

limit levels derived from mathematical analysis. 

House keeping telemetry monitoring is performed to know the satellite status. 

6.2.4.2 Random Vibrations 

Random vibrations test at satellite level is not foreseen. 

6.2.4.3 Acoustic Noise 

Acoustic noise qualification test will be performed on the integrated satellite. 

The expected levels seen by the payload will be covered by random qualification level specification of section 5.1.3. 

and acoustic qualification level specification of section 5.1.4. 

The test sequence will be as described hereafter: 

Low level, performed with 8 dB less than the qualification level, during 1 minute. 

Intermediate level, performed with 4 dB less than the qualification level, during 1 minute. 

Qualification level during 1 minute. 

Control low level, performed with 8 dB less than the qualification level during 1 minute to verify that vibration 

did not modified the behaviour of the Payload. 

House keeping telemetry monitoring is performed to know the satellite status. 

6.2.4.4 Pyrotechnic shocks 

Pyrotechnic tests will be performed on the integrated satellite simulating launch vehicle separation and satellite EEDs 

activation. 

The expected levels seen by the units will be covered by pyrotechnic shock qualification level specification of section 

5.1.5. 




6.2.5 EMC-Test 

6.2.5.1 Conducted emission 

• Electrical configuration is defined as below : 

EMC Conducted SL-EGSE Configuration 

Refer to Paylod Prime AIT requirements 

PL EGSE PLTM LAN 

PL EGSE 

SAS SADM to PCE 

SCOE PWR lines 

PL 

Umbilicals 

PCE 

Test Jig on PL pwr EMC CE 

& signal lines Acquisition 

equipment 

TTC Antenna TTC Test cap (isolated) on +Z 

GPS Antenna Test cap -X GPS 

(isolated) SCOE 

Aux Pwr RF 

SCOE SCOE 

PC for PLTM TM/TC 

ftp exchg. SCOE 

HK TM/TC & RC/RM LAN Ethernet 10 Mbps 

MCDT (SL bus EGSE) 

N.B.: Battery integration is required for this qualification as battery simulator EGSE is not reprentative of conducted 

EMC characteristics. 

• Conducted emission verification: 

A blank scan is performed to calibrate the background noise (test set up operating with satellite power switch 

off). 

Satellite is powered on to noisy mode : equipment in the CE worst cases 

Conducted emission is measured on BNR primary bus at PCE connection (ripple detection..) and at PF/PL I/F 

connector bracket 

6.2.5.2 Radiated emission and susceptibility 

These tests will be performed with the satellite in a shielded anechoic chamber on a tilting dolly MGSE, with a 

minimum hard-line connection to EGSEs. 

This test will be performed with the satellite in a shielded anechoic chamber on integration dolly MGS, with a 

minimum hard-line connection to EGSEs. 




• Electrical configuration is defined as below : 

EMC Radiated Configuration : Autocompatibility 

To Be Define for each Payload 

PL EGSE PLTM LAN 

PL EGSE 



PL 

EMC RE/RS 

field 

measur. 

equipment 

GPS 

Battery SCOE 

simulator 

PF 

TTC Antennas 

ANECHOIC Umbilical 

ROOM 

Aux Pwr RF 

SCOE SCOE 

PC for PLTM TM/TC 

ftp excgh. SCOE 

HK TM/TC & RC/RM LAN Ethernet 10 Mbps 

MCDT (SC bus EGSE) 

• SC Radiated EMC self-Compatibility: 

A blank scan is performed to calibrate the background noise (test set up operating with satellite power 

switched off). 

Platform reference measurement by field measurement and reference equipment HCT, 

Instrument characterisation alone and with PF equipment, 

Instrument characterisation all together and with PF equipment. 

• S/C EMC Radiated Emission and Susceptibility with launcher TBD: 

S/C removal from integration dolly and hanging on Hosting device by the room crane for EMC/RS 

susceptibility with launcher


R.E. : S/C is switched on in launch mode and radiated emission is performed within RF launcher required 

bands range (Radiated emission is measured at 1 m of I/F plane in circular polarisation) as specified in 

section 3.5.7.2.1. 

R.S.: an RF field is radiated according to launcher and launch site requirements, at 1 m by a test antenna , 

with a circular polarisation or two perpendicular axes with a linear antenna, as specified in section 

3.5.7.2.2. Susceptibility is checked by switching on the S/C to launch mode with required checks foreseen 

during combined operations. 

Satellite re-installation on Integration Dolly 

6.2.5.3 RF compatibility test 

Test sequence TBD. 

6.2.6 ESD-Test 

ESD test at satellite level is not foreseen. 





Chapter 7 : Generic PROTEUS control ground segment 








7. GENERIC PROTEUS CONTROL GROUND SEGMENT 5 

7.1 PURPOSE 5 

7.2 SCOPE 5 

7.3 PGGS FUNCTIONS 5 

7.3.1 STATION KEEPING PHASE 5 

7.3.2 FINAL ORBIT ACQUISITION PHASE 6 

7.4 OPERATIONS CONCEPTS AND OPERATIONAL ORGANIZATION 7 

7.5 ARCHITECTURE 8 

7.5.1 BASIC ARCHITECTURE 9 

7.5.2 ARCHITECTURE WITH OPTIONS 10 

7.5.3 SYSTEM TECHNICAL CHOICES FOR PGGS DEFINITION 11 

7.5.4 COMMUNICATION ARCHITECTURE 12 

7.6 INTERFACES BETWEEN PGGS COMPONENTS 13 

7.6.1 DIAGRAM AND LIST OF INTERFACES 13 

7.6.2 OPERATING MODES 15 

7.6.2.1 Telemetry processing operating mode 15 

7.6.2.2 Telecommand processing operating mode 15 

7.6.2.3 TTCET station management processing operating mode 16 

7.6.2.4 CCC-Mission Center interface operating mode 16 

7.6.2.5 Angular measurement and 2GHz KIT option operating mode 16 

7.7 PERFORMANCE 17 

7.7.1 PGGS MONOSATELLITE PERFORMANCE 17 

7.7.2 PGGS MULTISATELLITE PERFORMANCE 17 

7.8 CNES OPERATIONAL ORGANIZATION 18 

7.8.1 OPERATIONAL ORGANIZATION FOR STATION KEEPING 19 

7.8.2 OPERATIONAL ORGANIZATION FOR FINAL ORBIT ACQUISITION 19 


Figure 7.5-1 : Basic PGGS diagram........................................................................................................................ 9 

Figure 7.5-2 : Diagram of PGGS with options....................................................................................................... 10 

Figure 7.5-3 : Communication architecture........................................................................................................... 12 

Figure 7.6-1 : Interfaces between PGGS components ............................................................................................ 13 

Figure 7.8-1 : Relations between PGGS components and CNES multimission items ............................................... 18 





Table 7.6-1 : PGGS interfaces and their functions ................................................................................................. 14 





LIST OF TABLES...................................................................................................................................................... 3 


LIST OF TBCs ........................................................................................................................................................ 4 

LIST OF TBDs ......................................................................................................................................................... 4 



Sectio 

n 


LIST OF TBCs 

LIST OF TBDs 

Sentence Planned 

resolution 

§7.8.1 1. An organisation bases on the standard one but complemented by a hot line service 

TBD hours-a-day using: 




7. GENERIC PROTEUS CONTROL GROUND SEGMENT 

7.1 PURPOSE 

This chapter describes the architecture of the generic ground control segment for a satellite based on the PROTEUS 

platform, the operations concepts applicable to the various components, the data exchanges between the PGGS 

components and the data exchanges with other external components. 

This architecture is broken down into the final orbit acquisition and station keeping phases. 

7.2 SCOPE 

This part is applicable to all missions using the PROTEUS platform for all concerning satellite control-command. 

7.3 PGGS FUNCTIONS 

For a given mission, the PROTEUS generic ground segment (PGGS) is part of the Mission Ground Segment. It does 

not ensure all the functions of the mission but those required for satellite final orbit acquisition and station keeping. 

The PGGS functions are broken down according to the satellite life phase, the station keeping phase or the final orbit 

acquisition phase. 

The PGGS is multisatellite for a given mission. The multisatellite configuration is limited to a cluster of 3 to 4 

satellites. 

To fulfil these functions, the PGGS consists of 3 components: 

The Command Control Centre (CCC) 

The Telemetry and Telecommand Earth Terminal (TTCET) 

The Data Communication Network (DCN) 

7.3.1 STATION KEEPING PHASE 

- Satellite monitoring and technical control 

The satellite monitoring and technical control consists in checking by processing monitoring telemetry data 

(HKTM) that the state of the satellite meets mission requirements, in transmitting telecommands intended to 

maintain the normal operation of the satellite, in transmitting telecommands intended to obtain additional 

diagnosis data, in transmitting telecommands to rectify abnormal situations or to call on onboard redundancies. 

The technical abilities of the CCC cover the platform and payload aspects for all that which may endanger the 

satellite survival. 

- Satellite configuring 

Even though the PROTEUS family satellites are highly autonomous thanks to automatic control of the onboard 

flight software, a certain number of systematic operations must be performed on the satellite to maintain it in an 

operational mode and to optimise its life (calibration of sensors, measurements on tanks, etc.). The frequency of 

these operations is very low, typically every three months. Operations procedures are associated with all these 

operations. 

- Orbit and attitude controls 




The calculations are done onboard by the AOCS from GPS data, attitude sensor information, data catalogue and 

models (magnetic fields, star catalogue, etc.). It consists in updating the UT/atomic time delta ; the parameters of 

the reference system change model at a very low frequency (typically every month) in delivering the control and 

orbit control instructions to the AOCS at a specific frequency according to mission requirements. Operations 

procedures are associated with all these operations. 

- Payload service 

Payload service consists in transmitting to the payload processing centres (MC) the telemetry data produced by the 

payloads (PLTM) received on the ground via the TTCET, checking the state of the payloads thanks to monitor 

telemetry processing (HKTM) and performing the programming operations according to mission requirements. 

Programming frequency depends on mission requirements, the long duration onboard programming storage 

capability are used. Operations procedures are associated with each programming operation. 

- Satellite expert appraisal 

Satellite expert appraisal consists in leading investigations in case of anomalies, making out operations reports 

for experience feedback especially for PROTEUS platform changes. These investigations and reports are 

performed from archived monitoring telemetry data (HKTM) and various generated operational data (logbooks, 

telecommand logs, orbitography data). 

7.3.2 FINAL ORBIT ACQUISITION PHASE 

During this phase, the PGGS ensures the same functions as the ones performed during the station keeping phase : 

satellite monitoring and technical control 

satellite configuring 

orbit and attitude controls 

satellite expert appraisal. 

Except for the payload service, this function is not applied or reduced to the strict minimum defined specifically for 

each mission. 

However, the activities related to orbit and attitude controls are completed by manoeuvres intended to reach mission 

orbit. These calculations are optimised for each mission and result from the mission analysis. 

During this phase, the satellite is supposed to be, in relation to its nominal transfer orbit, inside the covariance matrix 

relevant to the launcher retained for the mission. Outside of this specification, maximum efforts must be made to 

recover the satellite without however being committed to a result. 




7.4 OPERATIONS CONCEPTS AND OPERATIONAL ORGANIZATION 

The operations concepts result from the general requirements and use the following platform characteristics : 

Requirements: 

To be able to perform the operations imposed by the platform during working hours. 

To be compatible with multiform mission ground segment organisations: CNES mission, mission in cooperation 

with other agencies, commercial or export missions. 

To be compatible with highly varied mission programming needs: once or twice per day, several times per 

month. 

To be modular to indifferently cover final orbit acquisition and station keeping needs. 

Platform characteristics: 

The platform is robust and able to protect itself from emergency case which occurs under 72 hours. This satellite 

autonomy avoids ground mechanisms which monitor satellite under 72 hours. 

The following concepts have been retained: 

The TTCET operates without operator and is controlled by the CCC. 

The CCC operates with operators working non-stop for final orbit acquisition. 

The CCC operates with operators working normal hours for station keeping. 

For final orbit acquisition and station keeping, presence of an operator at the CCC is required only for 

transmitting telecommands and preparing operation sequencing. 

The onboard and ground items can be monitored automatically from the CCC, an anomaly must call for 

operator intervention within a variable delay specific to each mission. 

All the CCC functions, other than telecommand transmissions, can be activated in automatic or manual 

sequencing mode. 

Generation of satellite commands is divided out and attributed to groups with independent responsibilities; 

there are three groups in all: the platform command generator group, the AOCS command generator group 

and the mission command generator group. 

Satellite expert appraisal is decentralised from the CCC; it can be performed directly by the expert from his 

work station. 

The state of the TTCET is controlled through the state of the TM data that it delivers; there is no need for 

station equipment monitoring. 

Satellite state is monitored in deferred time at a daily frequency varying according to mission requirements: 

from "after each pass" to "at least once every 7 days". 

A PGGS architecture and an operational organisation for final orbit acquisition and station keeping result from these 

operations concepts. The operational organisation presented is specific to CNES missions. 




7.5 ARCHITECTURE 

The PGGS is presented as a set of three basic components: the CCC, TTC-ET and DCN, to which various options are 

added, these options being: 

1. TTC-ET station duplication. 

2. Use of Angular Measurements obtained by the 2GHz stations for first acquisition. 

3. Use of a 2GHz station with PROTEUS TM/TC kit during the final orbit acquisition phase to increase 

operational availability (Option proposed within the scope of final orbit acquisition performed by CNES). 

4. Use of a 2GHz station with the PROTEUS TM/TC kit during routine phase in case of serious failure to the 

operational TTC-ET station (Option proposed within the scope of a CNES mission). 

Each component is based on the assembly of elementary building blocks. 

For the TTCET, the elementary building blocks are the following ones: 

1. "Antenna/Tracking" part 

2. "TM/TC Processing" part 

3. "Time Frequency" part 

Notice: part 2 can itself be broken down into two sub elements: the TM and the TC. 

For the CCC, the elementary building blocks are teh following ones: 

1. "Onboard/Ground" interface part 

2. "Orbit and Attitude" part 

3. "Archiving" part 

4. "Consultation/Expert Appraisal" part 

5. "Operations Automation" part 

The DCN consists in various networks supporting the IP protocol. 

The generic characteristic of the PGGS for the various missions is obtained by assembling components including all 

or part of their elementary building blocks and options. 

For example: 

The PROTEUS TM/TC kit added to a 2GHz station corresponds to the elementary building blocks 2 and 3 of the 

TTCET. 

For a single satellite mission, the CCC consists in 5 elementary building blocks; for a mission including several 

satellites, building blocks 1 and 3 are duplicated. 




7.5.1 BASIC ARCHITECTURE 

Figure 7.5-1 shows the basic PGGS. 

In this configuration, it meets the final orbit acquisition and station keeping requirements for a mission: 

Compatible with an operational availability, excluding launch, of 0.93. 

Multisatellites without need for simultaneous access to satellites. 

Use of a launcher with launch positioning errors in DELTA class. 

Figure 7.5-1 : Basic PGGS diagram 




7.5.2 ARCHITECTURE WITH OPTIONS 

Figure 7.5-2 shows the basic PGGS with its options. In this case, it meets the final orbit acquisition and station 

keeping requirements for a mission: 

With operational availability requirements, excluding launch, better than 0.93. 

Multisatellites. 

Using all types of launchers identified for PROTEUS 

Figure 7.5-2 : Diagram of PGGS with options 




7.5.3 SYSTEM TECHNICAL CHOICES FOR PGGS DEFINITION 

- Communications protocol 

The communications principle retained is "all IP" (UDP/IP or TCP/IP enabling file transfer (FTP), remote control (rlogin 

or telnet), remote execution (RPC) and real time transfer (socket stream or datagram). 

To obtain IP communication protocol on WAN, we will interconnect the local network (Ethernet) of the Control 

Command Center (CCC) to the local network (Ethernet) of the station (TTC-ET) by a standard router via the WAN 

communication means retained for the mission. 

The use of a standard router will enable us to more easily adapt to a later change in WAN communication means by 

simply replacing the router without modifying the CCC software or the TTC-ET. 

-Exchange of telemetry data in server customer mode between CCC and TTC-ET and between MC and TTC-ET 

In this mode, the TTC-ET, the recorded platform telemetry and payload telemetry data server and producer, places 

the data received, outside of the pass, at the disposal of the customers (CCC and MC). There is uncoupling between 

recorded platform telemetry or payload telemetry reception and its use by the customers. The customers take 

initiative for the exchange that they perform in accordance with their requirements. The real time telemetry is 

transmitted in real time to the CCC. 

- CCC-SAT telecommand link in authenticated mode 

The telecommands are segmented at the CCC and authenticated. Then all telecommand to onboard TC decoder are 

protected against intrusion. 

- End-to-end telecommand retransmission protocol (CCC-SAT) 

Transmission of telecommand segments is based on the COP1 protocol where the onboard automatism part is 

implemented in the onboard TC decoder, the ground automatism part is implemented in the CCC. Retransmission 

management covers the end-to-end link, from CCC work station to the onboard TC decoder output. 

- Open loop antenna designation 

The TTC-ET antenna is controlled by the designations calculated by the CCC. The simplification of the station 

equipment to comply with the cost reduction targets imposes this type of control mode. 

- Interface with mission center 

The PGGS receives the programming messages from a single mission center (MC-P). The programming messages 

are sets of ready-to-send TCs. Only segmentation and authentication are done by the PGGS. 

- Multisatellite design 

The PGGS software packages are established for a single satellite. Multisatellite implementation is achieved by 

duplicating the software on the same machine if operation is sequential or on different machines for simultaneous 

use. 




7.5.4 COMMUNICATION ARCHITECTURE 

The communication architecture is based on IP routers interfacable with the RNIC, the Integrated Services Digital 

Networks (ISDN), and dedicated lines (cf.Figure7.5-3). 

The security of the information systems can be ensured in three ways: 

by physical insulation of the IP network; this solution being well adapted to export use, 

by use of PACTE, 

or by use of a PGL. 

The PACTE solution is adapted to CNES missions. The PGL solution, specific to PROTEUS, is the solution to be 

retained if mission operational availability constraints are not compatible with those of PACTE. 

Data communication network (DCN) administration can be ensured in two ways: 

by station integrated into the CCC; this solution is adapted to export use, 

from one of the CNES network administration service stations for CNES missions. 

The IP address range retained for a given mission will belong to the CNES address ranges for CNES missions. 

Figure 7.5-3 : Communication architecture 




7.6 INTERFACES BETWEEN PGGS COMPONENTS 

7.6.1 DIAGRAM AND LIST OF INTERFACES 

The interfaces between the PGGS components are shown on Figure 7.6.1 and listed in details in Table 7.6-1. 

TTCET TTCET 

PLTM 

MC MC 

CCC - TTCET 

Telecommands (TC) 

Remote Controls (RC) 

Antenna designation 

TTCET - CCC 

HKTM-P, HKTM-R 

Remote Monitoring (RM) 

TC acknowledge 

RC acknowledge 

TTCET logbook 

Proc station - CCC 

CCC data request 

CCC - Proc station 

Consultable CCC data 

(HKTM-R, LogB, 

satellite status, etc) 

TTCET - MC 

CCC CCC 

Data Data remote remote processing processing PC PC 

CCC - MC 

Timing diagram 

TC transmission acknowledge 

Satellite orbit & attitude prediction 

MC - CCC 

Payload prog. TC 

CCC - Sband ET 

Telecommands 

Sband ET - CCC 

HKTM-P, HKTM-R, RM 

TC acknowledge 

Orbit parameters 

IERS data 

Solar activity 

Figure 7.6-1 : Interfaces between PGGS components 

2 GHz GHz 

Proteus Proteus Kit Kit 




GENERIC NAME ROLE 

CCC_MC_ORBIT_EVENTS Timing diagram giving all predicted events related to the 

orbit, to station visibilities or to AOCS programming for 

orbit and attitude control 

CCC_MC_PREDICTED_ATTITUDE Predicted satellite attitude [orbit] elaborated by the CCC 

for x days duration 

CCC_MC_PREDICTED_ORBIT Predicted satellite orbit elaborated by the CCC for x days 

duration 

CCC_MC_TC_LOGBOOK Report of Payload TCs and Platform TCs transmitted by 

CCC to satellite 

CCC_OCC_ORBIT_PARAMETERS Orbit parameters estimated by CCC and supplied to OCC (Orbit 

Computation Center) when recourse to CNES 2GHz network is 

required (acquisition on first orbits or survival). The OCC uses 

these parameters to calculate 2GHz network station designations 

CCC_TTCET_PASS-PLANNING Pass planning to be followed by TTCET sent by Main CCC 

CCC_TTCET_POINTING Antenna designations describing a pass of a visible satellite sent 

by the Main CCC to TTCET 

CCC_TTCET_RC All remote controls sent by Main CCC to TTCET 

CCC_TTCET_TC CLTU to CCSDS format containing the TCs sent by the Main CCC 

to TTCET 

MC_CCC_TCPL Payload TCs sent by MC to CCC for mission programming 

OCC_CCC_IERS_DATA Data delivered by IERS and transmitted to CCC via OCC 

OCC_CCC_ORBIT_PARAMETERS Orbit parameters delivered to CCC by OCC when recourse to 

CNES 2GHz network is required (acquisition on first orbits or 

survival) to locate the satellite (TTCET station designation accuracy 

insufficient). In this case, the OCC performs orbit determination 

from angular measurements and produces adjusted orbit 

parameters 

OCC_CCC_SOLAR_ACTIVITY Solar activity data delivered by the MEUDON observatory. 

Transmitted to CCC via OCC 

TTCET_CCC_ACQRC CCSDS Packets containing TTCET RC reception acknowledgments 

TTCET_CCC_CLCW CCSDS Packets containing the CLCWs extracted from TM frames 

and transmitted in real time by TTCET to Main CCC 

TTCET_CCC_HKTMP CCSDS Packets of real time telemetry transmitted in real time by 

TTCET to CCC (Main CCC and/or Secondary CCC) 

TTCET_CCC_HKTMR Files containing CCSDS packets of HKTM-R stored in TTCET 

TTCET_CCC_LOGBOOK Station logbook 

TTCET_CCC_RM CCSDS Packets containing TTCET remote monitoring information 

TTCET_MC_PLTM Files containing payload telemetry stored in TTCET 

Table 7.6-1 : PGGS interfaces and their functions 




7.6.2 OPERATING MODES 

7.6.2.1 Telemetry processing operating mode 

In TTCET 

TM reception at station 

Separation of TM flows 

Transmission of pass TM to CCC in real time (HKTM-P) 

Local storage of recorded TM (HKTM-R) in station and making available to CCC for 72 hours. 

Local storage of payload TM (PLTM) in station and making available to Mission Centers (MC) for 72 hours. 

In CCC 

Reception, demultiplexing for TC function and transmission of real time telemetry to DRPPC in real time 

Recovery of recorded plat-form telemetry after pass for archiving and processing (satellite monitoring, satellite 

state generation and orbit calculation). 

In MCs 

Specific mission processing operations 

7.6.2.2 Telecommand processing operating mode 

In CCC 

Elaboration of platform TCs and AOCS TCs. 

Recovery of ready-to-send mission TCs from MC. 

Authentication and Transmission of TCs to satellite with satellite current state checks. 

In main MC 

Elaboration of mission programming TCs 

In TTCET 

Setting up of onboard-ground link on CCC request. 

Transmission of TCs to satellite. 

Real-time transmission of TC acknowledgements received by satellite to CCC. 




7.6.2.3 TTCET station management processing operating mode 

In CCC 

Transmission of remote controls (RC) to TTCET to change TM digital rate, start and end of TC session, change 

of TC transmission polarization. 

Reception of RC acknowledgement after each transmission. 

Transmission of antenna designations and pass management at least every 72 h (12h for orbits at altitudes 

lower than 600 km). 

Transmission of a station long loop test for diagnosis in case of CCC-Satellite link failure. 

7.6.2.4 CCC-Mission Center interface operating mode 

From CCC 

Transmission of event timing diagram related to orbit, station visibilities and station programming after orbit 

determination. 

Transmission of TC transmission report after onboard loading. 

Recovery of MC mission TCs at mission frequency. 

7.6.2.5 Angular measurement and 2GHz KIT option operating mode 

At OCC for MA option 

Elaboration of orbit parameters from MAs 

Transmission of orbit parameters to CCC 

At CCC for 2GHz KIT option 

Elaboration of orbit parameters 

Transmission of orbit parameters to OCC 

At OCC 

Elaboration of station designation with KIT 




7.7 PERFORMANCE 

7.7.1 PGGS MONOSATELLITE PERFORMANCE 

Telemetry processing: 

Once set up, the end-to-end link ensures a maximum telemetry frame loss rate of 10 -8 

Recovery at TTC-ET of recorded TM (HKTM-R) over one day (40Mb) in less than 16 minutes during pass on 40 

Kbs channel. 

Access by CCC to recorded TM (HKTM-R) of one day, stored at TTCET in less than 15 minutes outside pass on 

a 64 Kbs link between CCC and TTC-ET. 

Telecommand processing: 

The end-to-end link, once set up, ensures a telecommand frame rejection probability lower than 10-5. 

The end-to-end link is protected by authentication guaranteeing probability of a successful attack (recognition 

of telecommand profile) lower than 10 -12 . 

Maximum rate is fixed by the capacity of the Satellite/TTC-ET link, that is 4Kbs. 

Designation: 

• The accuracy of TTC-ET designation, all causes combined (station and designation accuracy) enables 

complete autonomy of TTC-ET greater than 72H for all orbits with an altitude higher than 600 km, a 

minimum autonomy of 12H for orbits between 500 and 600 Km. 

• The accuracy of the designation data calculation (three-sigma angle) delivered by the CCC is 

guaranteed as better than: 

± 0.3 deg over 72H for orbits of altitude > 1000 km, 

± 0.5 deg over 72H for orbits of altitude between 800 and 1000 km, 

± 0.7 deg over 72H for orbits of altitude between 600 and 800 km, 

± 1 deg over 12H for orbits of altitude between 500 and 600 km. 

Operational availability: 

During the first 24 hours of final orbit acquisition PGGS availability is better than 0.97. 

Maximum repair time is 12 hours for simple failures with equipment in stock 

After the first 24 hours, PGGS availability is better than 0.93. 

Maximum repair time is 72 hours for simple failures with equipment in stock 

7.7.2 PGGS MULTISATELLITE PERFORMANCE 

Same performance is guaranteed for multisatellite configurations with following constraints: 

With a one TTC-ET configuration, two successive visibilities must be separated by at least 5 min. 

Access time by CCC to recorded telemetry (HKTM-R) of one day, stored in TTCET, will be complied with 

outside pass time of any one of the satellites. 




7.8 CNES OPERATIONAL ORGANIZATION 

Figure7.8-1 indicates the relations between PGGS components and CNES multimission items. 

-- 

TTCET 

Main CCC - TTCET 

TC (TeleCommand) 

RC (Remote Control) 

POINTING (Antenna pointing) 

PASS_PLANNING 

TTCET - Main CCC 

HKTMP, HKTMR 

RM (Remote Monitoring) 

CLCW (TC send 

Acknowledge) 

ACKRC (RC send 

Acknowledge) 

CCC - Ext users 

TC (TCs 

External 

Users 

CCC - MCR 

HKTMP 

RM 

CLCW 

TTCET - MC 

PLTM_FRAME 

PLTM PACKET 

Main CCC 

MC 

CCC - MC 

ORBIT_EVENTS 

TC_LOGBOOK 

PREDICTED_ORBIT 

PREDICTED_ATTITUDE 

MC - CCC 

TC_PL (Payload TCs 

SPC 

CCC - SPC 

RAWDUMP 

(Satellite Prime 

Contactor) 

SYMBOLICDUMP 

SPC - CCC 

MAPLV 

(Same IF descri ptions than TTCET - 

CCC - SBANDET 

TC, RC, PASS-PLANNING 

SBANDET - CCC 

HKTMP, HKTMR, RM, 

CLCW 

CCC - OCC 

ORBIT_PARAMETER 

S 

OCC - CCC 

ORBIT_PARAMETER 

S 

IERS_DATA 

SBANDET 

with Kit 

TTCET 

OCC 

DRPPC 

DRPPC - CCC 

CCC data request 

CCC - DRPPC 

Answers 

• HKTM 

DRPPC 

(CNES Clients, 

Main Control 

Room) 

• 

• 

• 

• 

• 

Logbooks 

Satellite status 

Email 

Satellite Data Base 

Visualisation files 

(CNES Clients, 

External 

Clients, Main 

Control Room) 

• Documentation 

Figure 7.8-1 : Relations between PGGS components and CNES multimission items 




7.8.1 OPERATIONAL ORGANIZATION FOR STATION KEEPING 

The platform design permits the following standard organisation for station keeping : 

an operator station which performs all daily operational tasks during working hours, 

a mission onboard-ground team (one or two engineers), 

Data Communications Network monitoring performed during working hours by the network department. 

The operator station is in charge of: 

preparing the task sequencer work plan, 

preparing platform telecommands, 

transmitting platform, AOCS telecommands and telecommands received by Mission Centre, 

monitoring Command Contrôle Centre data. 

The mission onboard-ground team is in charge of: 

analysing satellite operation, 

optimising operation, 

correction of anomalies. 

However for the mission availability, two other organisation types are conceivable: 

1. An organisation bases on the standard one but complemented by a hot line service TBD hours-a-day using: 

the remote call function, 

the Data Remote Processing Personal Computer. 

2. If the previous organisation types do not meet scientific, preoperational or operational mission requirements (for 

example Jason 1 is a mission with 24 hours-a-day operations), a specific organisation shall be defined according to 

the mission needs. 

7.8.2 OPERATIONAL ORGANIZATION FOR FINAL ORBIT ACQUISITION 

For final orbit acquisition, operations are ensured non-stop. 

The station keeping organisation is complemented by: 

a corresponding Operational Computation Center (OCC) and Network Operations Center to handle MA 

processing, 2GHz station designation with PROTEUS KIT if these options are retained for the mission, 

5 to 6 stations in Main Control Room for satellite experts. 





Chapter 8 : PROTEUS Generic Ground Segment (PGGS) 

Référence Fichier :PUM Ch08 22 10 04 du 22/10/04 

17:25 




Reproduction interdite © ALCATEL SPACE Company confidential 

Référence du modèle : M023-3



17:25 


8. PROTEUS GENERIC GROUND SEGMENT (PGGS) – MISSION CENTRE INTERFACES 4 

8.1 SUBJECT 4 

8.2 INTERFACES NOMENCLATURE 5 

8.3 CONVENTIONS APPLIED TO ASCII FILES 6 

8.4 PGGS – MISSION CENTRE INTERFACES DESCRIPTION 7 

8.5 NETWORK IF – FTP CONNECTIONS SPÉCIFICATIONS 39 

8.5.1 TRANSFER SCENARIO 39 

8.5.2 CONNECTION REQUIREMENTS 39 


Erreur! Aucune entrée de table d'illustration n'a été trouvée. 







LIST OF TABLES...................................................................................................................................................... 2 


LIST OF TBCs ........................................................................................................................................................ 3 

LIST OF TBDs ......................................................................................................................................................... 3 





17:25 

LIST OF TBCs 

LIST OF TBDs 




8. PROTEUS GENERIC GROUND SEGMENT (PGGS) – MISSION 

CENTRE INTERFACES 

8.1 SUBJECT 

The purpose of this chapter is to specify, for each exchanged data between the Mission Centre (MC) and PROTEUS 

Generic Ground Segment (PGGS), all the useful information for their understanding and treatment. 

The data are described using several different forms: 

FORM1 general level of interface description. 

FORM2 File general characteristics. 

FORM3 File logical records description giving their size, number and fields. 

FORM5 File example. 


17:25 




8.2 INTERFACES NOMENCLATURE 

Generic interface name XXX_YYY_FREE UPPER CASE LETTER TEXT 

Example: 

TTCET_CCC_HKTMP HKTM-P data provided by TTCET to CCC 

File name FREE TEXT WITH FOLLOWING RULES: 


17:25 

Separated character: _ (underscore) 

XXX Sender's abbreviation 

YYY Receiver's abbreviation with following rules: 

CCC Command Control Centre 

MC Mission Centre 

OCC Orbit Computation Center 

TTCET Telemetry TeleCommand Earth Terminal 

Separated character: _ (underscore) 

SLID is satellite identifier (if needed) 

ETID is earth terminal identifier (if needed) 

SLID is the satellite identifier defined with a character string (maximum 6 characters) 

(example JASON1 or COROT) 

ETID is a TTCET identifier defined with a character string (maximum 6 characters) 

(example AUS to design a CNES TTCET located at Aussaguel) 




8.3 CONVENTIONS APPLIED TO ASCII FILES 

C1 The lines which begin with the character # are comment lines 

C2 The first file record is #BEGIN_OF_FILE 

C3 The last file record is #END_OF_FILE 

C4 The real numbers are represented in scientific notation with a decimal point examples: 

1.2345E3 -1.2345E-3 123.456 -0.123456 

C5 The hexadecimal representation of values are terminated by the character h 

examples: A0F4h 042Ah 


17:25 




8.4 PGGS – MISSION CENTRE INTERFACES DESCRIPTION 


17:25 

GENERIC NAME ROLE 

CCC_MC_ORBIT_EVENTS Orbit sequence of events giving in anticipation orbital events, TTCET 

events (fly-by times) and satellite events (programming AOCS TC 

times) 

CCC_MC_PREDICTED_ATTITUDE Predicted satellite attitude information elaborated by the CCC to the 

MC 

CCC_MC_PREDICTED_ORBIT Predicted orbit data of the satellite and time reference (Position, 

Velocity, Time) elaborated by the CCC after an orbit determination 

CCC_MC_TC_LOGBOOK Sending acknowledge of TCPL and TCBUS transmitted from CCC to 

satellite 

MC_CCC_TC_PL Payload programming commands files provided by MC to CCC 

TTCET_MC_PLTM_FRAME Files containing PLTM CCSDS standard frames stored in TTCET and 

transmitted to MC on MC request 

TTCET_MC_PLTM_PACKET Files containing PLTM CCSDS standard packets stored in TTCET and 

transmitted to MC on MC request 




FORM1 INTERFACE DESCRIPTION FILE 

Generic interface name: CCC_MC_ORBIT_EVENTS 

Orbit sequence of events giving in anticipation orbital events, TTCET events (fly-by times) and satellite events 

(programming AOCS TC times) 

EXCHANGE DESCRIPTION 

Provider CCC Consumer MC 

Client CCC Server MC 

Protocol FTP authenticated mode Exchange initiative CCC 

Schedule Once a week and anytime if needed 

Comment 

EXCHANGE DATA DESCRIPTION 

Exchange format ASCII sequential file Compressed data NO 

File name SLID_ORBIT_EVENTS 

Size Max 1 Mbytes 


17:25 




File contains x records 

• Chronologically sorted file 

• 1 record contains 1 event description 

• The first record contains the file creation UT time 

• Event description parameters 

• Event time 

• Event class (Navigation, earth terminal, satellite or mission) 

• Event number in the class 

• Event orbital position 

• Event longitude and latitude 

• TTCET ID (only for earth terminal events class) 

• Event comment 

• List of events 

• Class of orbital events 

− Ascending and descending pass nodes times 

− Times of transitions (light -> half light -> shadow and reverse) 

− Times of under satellite point transitions (day -> night and reverse) 

− Time of shifting into quadrature position (satellite – sun – earth) 

− Time of shifting into subsolary position 

− Time of sun eclipse by moon 

• Class of Earth terminal events 

− TTCET AOS and LOS (0°, physical angle of elevation, any angle) 

− Maximal angle of elevation pass 

− TM/TC polarization modification 

− Times if TM/TC TTCET antenna glare by sun 

− RF AOS time and LOS time 

• Class of satellite events 

− AOCS TCs due date 

• Class of mission events (Mission dependent) 


17:25 




FORM2 FILE DESCRIPTION FORM 

FIlE NAME: SLID_ORBIT_EVENTS 

FILE DESCRIPTION 

Orbit sequence of events giving in anticipation orbital events, TTCET events (fly-by times) and satellite events 

(programming AOCS TC times) 

FILE TYPE 

Sequential [ X ] 

Number of record types: 2 

Logical structure of records: {«#BEGIN_OF_FILE», 

«1»,n*{«2»}, (n = number of events descriptions in the file) 

"#END_OF_FILE»} 

Direct [ ] 

Record size: 


17:25 

1 < UT file creation time> 

2 

 

NB: The lines which begin with the character # are comment lines 




FORM3 RECORD DESCRIPTION FORM 

FILE NAME: SLID_ORBIT_EVENTS 

Record number: 1 

Record size: 35 bytes 

RECORD DESCRIPTION 

Field name Size 

(bytes) 


17:25 

Kind Content description 

Line_Type 15 F ASCII Forced to 

File_Time 19 F ASCII File creation time 

(Format YYYY/MM/DD HH:MN:SS) 

All the fields are separated by a "tabulation character" 





FILE NAME: SLID_ORBIT_EVENTS 


Record size: Max 313 bytes 



(bytes) 

Event_Time 23 F ASCII Event time 


17:25 


(Format YYYY/MM/DD HH:MN:SS.MMM) 

Event_Class 1 F ASCII Class of Event 

O Orbital 

E Earth terminal 

S Satellite 

M Mission 

Class of NON SELECTED Event 

XO Non selected Orbital event 

XE Non selected Earth terminal event 

XS Non selected Satellite event 

XM Non selected Mission event 





(bytes) 


17:25 


Event_Number 2 V Int16 Event number in the class 

• Orbital events class 

1 Ascending node pass time 

2 Descending node pass time 

3 Light penombra transition time 

4 Penombra shadow transition time 

5 Shadow penombra transition time 

6 Penombra light transition time 

7 Day night transition time 

8 Night day transition time 

9 Time of shifting into quadrature position (satellite – 

sun – earth) 

10 Time of shifting into subsolary position 

11 Time of sun eclipse by moon 

• Earth terminal events class 

1 0° TTCET AOS time 

2 Physical TTCET AOS time 

3 Another fixed TTCET AOS time (5° for example) 

4 Another fixed TTCET LOS time (5° for example) 

5 Physical TTCET LOS time 

6 0° TTCET AOS time 

7 Maximum angle of elevation pass time 

8 Left right TM/TC polarization modification 

9 Right left TM/TC polarization modification 

10 Start of TM/TC TTCET antenna glare by sun 

11 End of TM/TC TTCET antenna glare by sun 

12 RF AOS time 

13 RF LOS time 

• Satellite events class 

1 Orbital position guidance TC 

2 Profile guidance TC 

3 SADM guidance TC 

4 Request STAM1 mode TC 

5 Request STAM2 mode TC 

6 Request OCM2 mode TC 

7 Beginning of thrust in OCM2 mode 

8 End of thrust in OCM2 mode 

9 Request OCM4 mode TC 

10 Beginning of thrust in OCM4 mode 

11 End of thrust in OCM4 mode 

12 Kinetic momentum TC 

13 Enable star tracker TC 

14 Disable star tracker TC 

15 Manoeuver beginning 

16 Manoeuver end 

• Satellite events class 

Mission dependent 





(bytes) 

Orbital_Position 6 F Real 

F6.2 

Longitude 6 F Real 

F6.2 

Latitude 6 F Real 

F6.2 


17:25 


Orbital position of the event 

Angle in degree from 0 deg (Equator) to 360 deg in the orbit 

direction 

Terrestrial longitude of the event 

Angle in degree from 0 deg (Greenwich meridian) to 360 deg in 

the East direction 

Latitude of the event 

Angle in degree from 0 deg (Equator) to +90 deg (North pole) and 

from 0 deg (Equator) to –90 deg (South pole) 

ETID 6 V ASCII Earth terminal identifier (only for the Earth terminal class, nothing 

otherwise) 

Comment 256 V ASCII String of characters describing the event 





FORM1 INTERFACE DESCRIPTION FORM 

Generic interface name: CCC_MC_PREDICTED_ATTITUDE 

Predicted satellite attitude information elaborated by the SOCC AOCS subsystem to the MOCC 




Protocol FTP authentified mode Exchange initiative CCC 

Schedule Depending mission requirements 

Comment • Covered period : depending mission requirements 

EXCHANGED DATA DESCRIPTION 


17:25 

• The first point is dated at the end of adjustment period 

• Fixed gap of 60 s between each point 

• If needed, the file takes a maneuver into account 


File name SLID_PREDICTED_ATTITUDE 

Size Variable 

• File contains x records 

• The first record contains the file creation UT time and the type of reference frame (J2000, WGS84 

or other) 

• Fixed length records 

• Record structure 

• UTC time of attitude event 

• Quaternion of the predicted attitude 

• Satellite attitude in ROLL, pitch and yaw (rd) 

• 3 components of the predicted satellite rate (rd/s) 

• Predicted position for the SADM 





FILE NAME: SLID_PREDICTED_ATTITUDE 


Predicted satellite attitude information elaborated by the CCC AOCS subsystem to Mission Center 

FILE TYPE 



Logical structure of records: {“#BEGIN_OF_FILE”, 

“1”,n*{“2”}, (n = number of points in the file) 

#END_OF_FILE”} 

Direct [ ] 

Record size: 


17:25 

1 

2 < QISLPRED2> < QISLPRED3> 

 

< SLRATEPREDY> < SLRATEPREDZ> 

 

NB: The lines which begins with the character # are comment lines 










(bytes) 

Line_Type 1 F ASCII Forced to 1 


17:25 




Type_frame 1 F Integer Type of reference frame in the file 

1 = J2000 

2 = WGS84 








Record size: Variable 



(bytes) 



17:25 


UTC_time 23 F ASCII UTC time of the attitude data 


QISLPRED1 V Float32 Component 1 of the predicted satellite attitude 




ROLLPRED V Float32 Predicted roll unit: rd 

PITCHPRED V Float32 Predicted pitch unit: rd 

YAWPRED V Float32 Predicted yaw unit: rd 

SLRATEX V Float32 Component Xs of the predicted satellite rate unit: rd/s 

SLRATEY V Float32 Component Ys of the predicted satellite rate unit: rd/s 

SLRATEZ V Float32 Component Zs of the predicted satellite rate unit: rd/s 

POSPREDL V Float32 Predicted position for the left SADM unit: rd 

POSPREDR V Float32 Predicted position for the left SADM unit: rd 






Generic interface name: CCC_MC_PREDICTED_ORBIT 

Predicted orbit data of the satellite and time (Position, Velocity, Time) elaborated by the CCC after an orbit 

determination 



Client MC Server CCC 

Protocol FTP authentified mode Exchange initiative CCC 

Schedule Depending mission requirements 

Comment 



Files name PREDICTED_ORBIT_SLID 

Size Variable 


17:25 

• File contains x records 

• The first record contains the file creation UT time and the type of reference frame (J2000, WGS84 

or other) 

• Fixed length records 

• Record structure 

• UTC time of orbit data (Position, Velocity) 

• Position (x, y, z) (m) 

• Velocity (vx, vy, vz) (m/s) 





FILE NAME: PREDICTED_ORBIT_SLID 


Predicted orbit data of the satellite and time (Position, Velocity, Time) elaborated by the CCC after an orbit 

determination 

FILE TYPE 




“1”,n*{“2”}, (n = number of points in the file) 

#END_OF_FILE”} 

Direct [ ] 

Record size: 


17:25 

1 

2 











(bytes) 



17:25 




Type_frame 1 F Integer Type of reference frame in the file 

1 = J2000 

2 = WGS84 











(bytes) 



17:25 


UTC_time 23 F ASCII UTC time of the orbit data 


X_Position 20 F Real 

F20.5 

Y_Position 20 F Real 

F20.5 

Z_Position 20 F Real 

F20.5 

X_Velocity 20 F Real 

F20.5 

Y_Velocity 20 F Real 

F20.5 

Z_Velocity 20 F Real 

F20.5 

X position (m) 

Y position (m) 

Z position (m) 

X velocity position (m/s) 

Y velocity position (m/s) 

Z velocity position (m/s) 






Generic interface name: CCC_MC_TC_LOGBOOK 

Sending acknowledge of TCPL and TCBUS transmitted from CCC to satellite 




Protocol FTP authenticated mode Exchange initiative CCC 

Schedule Depending on mission requirements 

Comment 



File name R_TCLOG_SLID_(YYYY_MM_DD_HH_MM_SS) begin__(YYYY_MM_DD_HH_MM_SS) end 

(extraction period UT date) 

Size Depending on number of sending TC during the period 


17:25 

• The first record contains the sending time of the first TC described in the file 

• The second record contains the sending time of the last TC described in the file 

• The following blocks describe the TCs 

• Block description 

• TC description (mnemo, destination, nature, operational description, APID, family, TC ID, 

MAP number, VC number) 

• TC sending time 

• Due date for time-tagged TC 

• TC sending acknowledge result 

• TC binary profile 





FILE NAME: R_TCLOG_SLID_YYYY_MM_DD_HH_MM_SS 


Sending acknowledge of TCPL and TCBUS transmitted from CCC to satellite 

FILE TYPE 



Logical structure of records: {«#BEGIN_OF_FILE», 

«1»,»2»,n*{3}, 

#END_OF_FILE»} 

(n = number of TC logbook messages) 

Direct [ ] 

Record size: 


17:25 

1 

2 

3 






FILE NAME: SLID_TC_LOGBOOK 





(bytes) 


TC_Mnemo 8 V ASCII TC mnemo 

TC_Dest 4 V ASCII TC destination 

TC_Nature 2 F ASCII TC nature 


17:25 


TC_OpDesc 80 V ASCII TC operational description 

TC_APID Int16 TC packet APID number 

TC_Family Int16 TC family number (0 for TCD) 

TC_ID Int16 TC number (0 for TCD) 

MAP Int16 Multiplexed access point number 

VC_ID Int16 Virtual channel number 

Sending_Time 19 F ASCII Sending time of the TC 


Due_Date 23 F ASCII Due date for time tagged TC 


TC_ACK 3 V ASCII TC acknowledge by satellite through the CLCW 

TC_Binary HEXA Binary TC profile in hexadecimal 

OK TC sent by CCC and acknowledged by the satellite 

NOK TC sent by CCC and non acknowledged by the 

satellite 






Generic interface name: MC_CCC_TC_PL 

Payload programming commands files provided by MC to CCC 


Provider MC Consumer CCC 


Protocol FTP authentified mode Connection initiative CCC 

Schedule Depending on mission requirments 

Comment 



File name SLID_TC_specific-name_YYYY_MM_DD_HH_MM_SS (File creation UT time) 

Size Max: 500 Kbytes 


17:25 

• 1 file contains ASCII description and binary profile TC 

• The first record contains the file creation UT time 

• The second record contains the provider of the file 

• 1 file contains one or more blocks of TC description 

• Each TC is described in a block which contains 

• TC mnemo and operational description 

• Due date if TC time-tagged 

• Delay before the TC sending 

• ASCII TC parameters description 

• Binary TC packet profile 

• For a time_tagged TC, it shall not specify a delay before the TC sending 





FILE NAME: SLID _TC_specific-name_YYYY_MM_DD_HH_MM_SS 


Payload programming commands files provided by MC to CCC 

FILE TYPE 




“1”,”2”,n*{“3”,[“4”],[”5”],[ m*{“6”}],”7”}, 

“#END_OF_FILE”} 

(n = number of TC description in the file) 

(m = number of ASCII TC description record) 

1 

2 

3 

4 (optional record) 



7 

Direct [ ] 

Record size: 


17:25 









RECORD DESCRIPTIO 

Field name Size (bytes) Kind Content description 

Line_type 15 F ASCII Forced to 

Creation_Time 19 F ASCII File creation UT time (Format YYYY/MM/DD HH:MN:SS) 


17:25 






FILE NAME: 





17:25 

SLID _TC_specific-name_YYYY_MM_DD_HH_MM_SS 



Provider 4 F ASCII TC group provider acronym 












TC_Mnemo 11 V ASCII Satellite Data Base TC mnemo 

TC_OpDesc 80 V ASCII Satellite Data Base TC operational description 




Record number: 4 Optional record 





Due_Date 

23 

F ASCII Due time for TC time-tagged 



17:25 












Delay V Int32 Delay to respect before the TC sending in milliseconds 

This field is authorized only if the TC is not a time_tagged command 









TC_Desc 80 V ASCII Free text describing the TC data 


17:25 











Line_type 12 F ASCII Forced to < TC_PROFILE> 

Length V Int16 TC packet length in bytes (max 248 bytes) 

TC_Binary V Hexa Binary TC packet profile in hexadecimal 

See packet structure in [RD2] 

• Packet header (6 bytes) with APID and data length 

• Packet data (max 242 bytes) 


17:25 






Generic interface name: TTCET_MC_PLTM_FRAME 

Files containing PLTM CCSDS standard frames strored in TTCET and transmitted to MC on MC request 


Provider TTCET Consumer MC 

Client MC Server TTCET 

Protocol FTP authenticated mode Exchange initiative MC 

Schedule Files creation after each programmed fly-by 

Comment • Data are provided by TTCET to MC after fly-by LOS + 3 min 



17:25 

• Storage time at TTCET level: 72 hours 

• 1 file contains an integer number of PLTM CCSDS frames 

• The PLTM file can be removed in TTCET by the MC after recovery and processing 

Exchange format Binary sequential file Compressed data NO 

File name SLID_PLTM1_F_YYYY_MM_DD_HH_MM_SS (If VC for PLTM1) 

SLID_PLTM2_F_YYYY_MM_DD_HH_MM_SS (If VC for PLTM2) 

(File creation UT time) 


• 1 file contains maximum 8900 frames of PLTM1 or maximum 8900 frames of PLTM2 

• Fixed-length records 

• 1 record contains 1 PLTM1 frames: 

• Frame synchronization marker (4 bytes) 

• CCSDS main header of the frame (6 bytes) 

• Data zone of the frame (1105 bytes) with: 

• Count of the virtual channel frame extension 

• PLTM1 or PLTM2 packet(s) 

• Operational control zone (4 bytes) 





FILE NAME: SLID_PLTMi_F_YYYY_MM_DD_HH_MM_SS 


Files containing PLTM CCSDS standard frames strored in TTCET and transmitted to MC on MC request 

FILE TYPE 



Logical structure of records: {n*{«1»}} (n = number of PLTM frames in the file) 

1 

 

Direct [ ] 

Record size: 


17:25 





FILE NAME: SLID_PLTMi_F_YYYY_MM_DD_HH_MM_SS 





(bytes) 


17:25 


Synchro 4 F Frame synchronization marker 

Forced to value 1ACFFC1D hexa 

Frame_Header 6 F Main header of the frame 

See frame structure in [RD2] 

Frame_Data 1105 F Data zone of the frame 


Control_Zone 4 F Operational control zone 






Generic interface name: TTCET_MC_PLTM_PACKET 

Files containing PLTM CCSDS standard packets stored in TTCET and transmitted to MC on MC request 


Provider TTCET Consumer MC 

Client MC Server TTCET 

Protocol FTP authenticated mode Exchange initiative MC 

Schedule Files creation after each programmed fly-by 

Comment - Data are provided by TTCET to MC after fly-by LOS + 5 MN 

- Storage time are TTCET level: 72 hours 

- 1 file contains an integer number of PLTM CCSDS packets 

- The PLTM file can be removed in TTCET by the MC after recovery and processing 


Exchange format Binary sequential file Compressed data 

File name SLID_PLTM1_P_YYYY_MM_DD_HH_MM_SS (If VC for PLTM1) 

SLID_PLTM2_P_YYYY_MM_DD_HH_MM_SS (If VC for PLTM2) 

(File creation UT time) 



17:25 

• 1 file contains x MN of PLTM1 or x MN of PLTM2 

• Variable-length records 

• 1 record contains 1 PLTM packet: 

• CCSDS packet header with in particular: 

• TM APID number 

• Data length (fixed length for each APID) 

• PLTM packet data (max 1018 bytes) with: 

• UT time except on-board computer time during safe mode 

• APID data (see Satellite Data Base) 





FILE NAME: SLID_PLTMi_P_YYYY_MM_DD_HH_MM_SS 


Files containing PLTM CCSDS standard packets stored in TTCET and transmitted to MC on MC request 

FILE TYPE 



Logical structure of records: {n*{«1»}} (n = number of PLTM packets in the file) 

1 

Direct [ ] 

Record size: 


17:25 





FILE NAME: SLID_PLTMi_P_YYYY_MM_DD_HH_MM_SS 


Record size: Max 1 kbytes 



(bytes) 


17:25 


Packet_Header 6 F Packet header including APID number and data length 

See TM packet structure in [RD2] 

Packet_Data V TM packet data with: 

• Datation information (10 bytes) 

• PLTM TM data (max 1008 bytes) 

See TM packet structure in [RD2] 




8.5 NETWORK IF – FTP CONNECTIONS SPÉCIFICATIONS 

8.5.1 TRANSFER SCENARIO 

CCC always initiates the FTP connections with TTCET, MC or OCC (i.e. CCC host is always client). 

MC always initiates the FTP connection s with TTCET. 

8.5.2 CONNECTION REQUIREMENTS 

R1 The FTP connection must be authenticated. 

R2 Login/password are included in system file "passwd". 

R3 A password is not in plain text. 

R4 The password must be changed at least every ninety days. 

R5 Files must be stored in a dedicated data directory. 

R6 The FTP server must provide a logfile. The server administrator only manages this logfile. 


17:25 





Chapter 9 : On board ground interface 




Status 

Reason of Change Change Reference Doc Issue 

§9 New in Useful TM data rate PUM.6.1.CG.06 6.2 







No Figure in this chapter 


Table 9-1: Frequency couples reserved at UIT for PROTEUS .................................................................................... 4 



LIST OF TBCs 

LIST OF TBDs 




Chapter 9: On board-ground interfaces 

The on board-ground interfaces designate the communication links between the ground control/command station or 

stations and the platform. It includes the interface with the launch and tests facilities. 

Trough this ground to satellite interfaces : 

communication can be established between the satellite and ground control station as well as the mission 

station(s) and can be maintained during visibility phases, 

communication can be protected against perturbations in order to achieve a minimum bit rate error, 

information can be transmitted from the platform and payload via the control station, thus ensuring that the 

status and functioning of the satellite can be monitored, 

commands can be transmitted from the ground to the platform and payload via the control station, 

data necessary to fix the satellite’s orbit can be exchanged with the ground and with the on-board equipment, 

communication with the satellite subsystems is possible, thus ensuring commands to the subsystem and 

acquisition of test results during integration, 

information can be transmitted between the payload and the ground control station and/or the mission 

station(s) for transfer to the mission user. 

If the User needs to have more information on the on board-ground interfaces, he can contact CNES or ALCATEL 

SPACE to get the interesting part or the whole of the document « Technical requirements specification : satellite-toground 

interface » (LDP-SB-LB/LS-12-CNES, issue 5). 

This document describes in details the on board-ground interfaces specifications for PROTEUS based missions. It 

deals with the following main subjects : 

1. The information flows which include 

the different telecommands types, 

the telemetry divided in permanent housekeeping telemetry (HKTM-P), housekeeping telemetry historic 

(HKTM-R), the failure diagnostic telemetry (FDTM) giving an accurate telemetry over a short period preceding 

a platform failure, the payload telemetry (PLTM1 and PLTM2), 

the separation flows and priorities, 

the volumes of the data flows. 

These aspects are already dealt in the PROTEUS User's Manual chapter 3. 

2. The exchange format which explains in details: 

the parameter number in a transmitted flow, 

the telemetry flow structure compliant with the ESA packet telemetry standard, the packet structure, the frame 

structure, the exchange format at the parameters level, 

the telecommand flow structure compliant with the ESA packets telecommand standard, the packet structure, 

the segment structure, the transfer frame structure, the transmission-units structure. 

3. The exchange protocols; that means the TM and TC circuits, the establishment and maintenance of the 

satellite to station link. 

4. The radioelectric interface 

The main characteristics are summarised below, and more accurate information is contained in the document « 

Technical requirements specification : satellite-to-ground interface » (LDP-SB-LB/LS-12-CNES, issue 5). 




The PROTEUS satellite - Ground link is ensured by a S band TM/TC link supporting CCSDS packet ESA standard 

data flow. 

The source packet shall not exceed a 512*16 bits word size (1 kbyte). 

The useful up link (TC) rate is equal to 4 kbit/s during all the satellite lifetime. 

The down link (TM) rate corresponds to 85.966 kbit/s [useful bit rate before coding] during the emergency phase 

and to 722.116 kbit/s during the routine phase. 

TM reception can be performed in both right and left circular polarisations. 

In Earth pointing mode, left polarisation will be used for TC in normal mode. 

For inertial, or solar pointing, the ground polarisation is modified depending on the current attitude during each 

visibility. 

For each mission, the telecommunication frequencies are chosen among the following frequency couples presented 

hereafter. The frequency couple chosen for a satellite is determined by both the customer and the PROTEUS team 

depending on the orbit and frequencies already attributed. If needed, other frequencies can be requested. 

Up link frequency Down link frequency UIT publication 

Couple 1 2040.34300 - 2040.64300 MHz 2214.920 - 2216.920 MHz AR/11A/1828 



Table 9-1: Frequency couples reserved at UIT for PROTEUS 

5. the exchange constraints which shall be respected by the ground and by the on board software are listed. 





Chapter 10 : Standard schedule, deliveries and documentation 




Status 


Reference 

§10.3.5 Modified in Star Tracker Assembly CIIS.4.1.JC.1_14 6.2 

§10.3.6.1 [PL - 10.3.6 -1 a] Modified in Connectors provided by 

ALCATEL 

CIIS.4.1.JC.2_3 6.2 

§10.3.6.2 Modified in Wiring provided by ALCATEL PUM.6.1.CG.31_31 6.2 

§10.3.6.2 [PL - 10.3.6 -2 a] Modified in Connectors provided by Payload CIIS.4.1.JC.2_3 6.2 




Doc Issue



10.1 PROTEUS STANDARD SCHEDULE (PRELIMINARY) 4 

10.1.1 PROTEUS PRE-PHASE B 7 

10.1.2 PROTEUS PHASE B 8 

10.1.3 PROTEUS PHASE C 9 

10.1.4 PROTEUS PHASE D 10 

10.1.5 PROTEUS PHASE E 10 

10.2 SATELLITE DOCUMENTATION 11 

10.2.1 APPLICABLE DOCUMENTS AND INTERFACES DOCUMENTS 11 

10.2.1.1 Satellite and system 11 

10.2.1.2 Launch vehicle interfaces 11 

10.2.1.3 Payload interfaces 12 

10.2.2 MANAGEMENT DOCUMENTS 14 

10.2.3 PRODUCT ASSURANCE DOCUMENTS 14 

10.2.4 DOCUMENTS RELATING TO DEVELOPMENT AND VALIDATION LOGIC 15 

10.2.4.1 Development and Validation plans 15 

10.2.4.2 Specifications of facilities for satellite integration and system tests 15 

10.2.5 TEST PLANS, ASSEMBLY INTEGRATION AND TESTS 16 

10.2.6 SYSTEM DESCRIPTION AND PERFORMANCES 16 

10.2.7 JUSTIFICATION DOCUMENTS 17 

10.2.8 SYSTEM DATA BASE AND OPERATIONS 17 

10.2.9 MISSION CENTRE/PROTEUS GENERIC GROUND SEGMENT INTERFACES DOCUMENTATION 17 

10.3 PROTEUS PROVISION 18 

10.3.1 PAYLOAD MECHANICAL MATHEMATICAL MODELS 18 

10.3.2 PAYLOAD CAD MODELS 18 

10.3.3 PAYLOAD FUNCTIONAL INTERFACES MODEL (EM OR PAYLOAD SIMULATOR ?) 18 

10.3.4 DELETED 18 

10.3.5 STAR TRACKER PROVISION 18 

10.3.6 PLATFORM/PAYLOAD INTERFACES WIRING PROVISION 18 

10.3.6.1 Platform/Payload power interfaces 18 

10.3.6.2 Nominal & Redundant Platform/Payload TM/TC interfaces 19 

10.3.7 THERMAL MLI 19 

10.3.8 PLATFORM/PAYLOAD INTERFACES SCREWS 19 

10.3.8.1 Payload interface screws 19 

10.3.8.2 STA interface screws 19 

10.3.8.3 Electrical brackets interface screws 19 







Figure 10.1-1 : PROTEUS standard schedule .......................................................................................................... 5 

Figure 10.1-2 : PROTEUS development logic .......................................................................................................... 6 


CHANGE TRACEABILITY Chapter 10 ...................................................................................................................... 1 

LIST OF TBCs 

LIST OF TBDs 




Chapter 10: Standard schedule, deliveries and documentation 

10.1 PROTEUS STANDARD SCHEDULE (PRELIMINARY) 

The generic schedule corresponding to the PROTEUS standard services is presented Figure 10.1-1. 

This schedule is built with the following hypothesis : 

• the satellite is based on a standard platform; the platform adaptations are limited to minor changes so called 

«missionisation». 

• ALCATEL SPACE and CNES lead activities on platform and satellite engineering, integration and tests. 

• The generic ground control segment is procured including one ground station and one control centre. 

• A single interface is considered between the mission centre and the control centre located in Toulouse. 

• PROTEUS standard services include the transportation, the launch campaign activities, flight acceptance & first 

operations and the control centre operations too. 

• Pre-Phase B and phase B durations are indicative. They shall be adapted to cope with payload development. 

Notice : BV is validation bench and may be adapted either in numerical validation bench or in system functional 

validation bench ; it is defined in the paragraph 10.1.3. 

Note that the satellite is based on an existing platform and consequently the schedule critical path is more likely 

through the payload development. Thus, payload works have to start firstly, but some satellite studies must begin as 

well, at the beginning of the payload development, in order to ensure a global technical consistency and provide 

payload with updated interface and environmental specifications. 

The logigram shown in Figure 10.1-2 describes the general logic regarding the satellite segment level and the 3 

main system levels which are Payload, Platform and Satellite On Board SoftWare. 



.. 


Figure 10.1-1 : PROTEUS standard schedule 




PLATFORM SATELLITE OBSW PAYLOAD 

PF 

Long 

Lead 

Standard 

elements 

PF 

AIT 

PF DRB 

PF 

mission 

adaptable 

elements 

design and 

manufact. 

Sat 

Spec. 

PUM 

Pre B Sat. 

SAT. 

B Phase 

SAT. PDR 

SAT. 

C Phase 

SAT. 

CDR 

SAT. 

AIT 

QFAR 

Launch 

Campaign 

Launch 

BV 

Adapt. 

SFC 

Validation 

PL DP0 

PDIS1 

PL DP1 

PDIS 2 

PL DP2 

PDIS 3 

PL EFM 

OBSW 

Adapt. 

OBSW 

Validation 

PL 

C Phase 

PL Perf. 

Spec. 

PL 

A Phase 

PL SRR 

PL 

B Phase 

PL PDR 

PL CDR 

PL AIT 

PL 

QR/DRB 



SWQR 

Figure 10.1-2 : PROTEUS development logic 

PL 

EFM 

Manufactur.


The satellite schedule is divided in five main phases : Pre-phase B, phase B, phase C (duration of 9 months), phase 

D (14 months)and phase E. Hereafter are described the main activities and objectives for each satellite phase. These 

essential points shall be fulfilled to maintain the schedule. 

10.1.1 PROTEUS PRE-PHASE B 

The purpose of this phase is to confirm that the mission belongs to the PROTEUS flight domain (feasibility study) and 

to provide the payload with specific inputs to complete the generic ones which are described in the PROTEUS User’s 

Manual (PUM). In order to achieve this goal, the necessary inputs, required at the beginning of this phase, are : 

• a first issue of Satellite Specification 

• a first issue of Mission Specification 

• a first Payload Data Package (DP 0) containing at least : 

• an issue of Payload Interfaces Requirements Descriptions or Instruments Interfaces Requirements Descriptions 

(if the payload is composed of several independent instruments and managed separately) 

• Payload mathematical models 

A simplified CAD model 

A simplified Finite Element Model 

• Payload budgets 

At the beginning, ALCATEL SPACE reads and comments the Satellite Specification, as well as the activities led by the 

Customer during the payload phase A. Then, the feasibility studies (depending on the satellite specific points) are 

performed and end at a Baseline Design Review (BDR) with : 

• A preliminary concept of the satellite, 

• An identification of points out of PROTEUS flight domain, 

• An identification of critical points, 

• A confirmation of schedule aspects. 

• A first issue of Payload Design and Interface Specification (PDIS) 

The Customer ensured the follow up of payload phases A and B. The input data provided by the platform for these 

first payload feasibility and design phases are : 

• the PROTEUS User’s Manual 

• the standard Star Trackers Assembly Finite Element Model (this standard STA can be modified if imposed by 

mission constraints; in this case, a new FEM will be provided at the beginning of payload phase C) 

The PUM is replaced by the PDIS at the end of this pre-phase B. During this phase, the customer still ensures the 

payload follow up. As an option, ALCATEL SPACE could participate to the payload meetings if ALCATEL SPACE is in 

charge of the payload follow up for the next phases. 




10.1.2 PROTEUS PHASE B 

Phase B starting defines the T0. 

The platform equipment are ordered during this phase and the long lead Items procurement starts immediately after 

T0. Some agreements between Alcatel Space and the concerned suppliers permit to get these items on time, that 

means for platform assembly. 

The objectives consist in establishing a satellite preliminary definition, in confirming the missionisation activities 

(platform harness, system data base, software updates and so on) to lead in phase C and also to provide the 

payload with accurate data for its detailed definition phase. 

In order to manage the phase B tight schedule and to obtain results as accurate as possible, the necessary inputs at 

the beginning of this phase are : 

• The launch vehicle choice. This point is very important to make accurate satellite qualification and flight 

environment requirements. 

• Update (if any) of the Satellite and Mission specification 

• A new payload Data Package (DP 1) containing at least : 

• A Payload Description Document (synthesis document) 

• A Payload Interface Control Document 

• A set of mathematical models 

• A CAD model 

• Two Finite Element Models (one physical model and one reduced model) 


• A Payload Design, Development and Validation Plan 

The required documents are described in section 10.2. 

Based on these data, the following main activities are performed : 

• Preliminary analyses in mechanical, thermal, electrical, Attitude Orbit Control System (AOCS) and command 

control domains. 

• Attitude Orbit Control laws coefficients tuning 

The Satellite Preliminary Design Review concludes the phase B and allows to begin the detailed analysis. At the end 

of this phase: 

• The payload interfaces specifications are updated in the PDIS (Payload Design and Interfaces Specifications). 

As an option, the instruments interfaces specifications are defined too in the IIS (Instruments Interfaces 

Specifications) and the PDIS is adapted at instrument level if ALCATEL SPACE is responsible for instruments 

integration in the Payload Instruments Module. 

• The ground and launch interfaces are defined too. 

• A satellite configuration is defined 

• The PL ICD is commented and, if necessary, STA and H02 & H03 new accommodation is proposed. 




10.1.3 PROTEUS PHASE C 

During this phase, Alcatel Space leads all the activities relating to the missionisation : 

• satellite detailed analysis in mechanical, thermal, electrical, Attitude Orbit Control System (AOCS) and 

command control domains. 

• platform realization files updates 

• flight software updates 

• satellite data base parameters updates. 

• Payload Design and Interfaces Specifications (PDIS) are updated again after detailed analysis. 

• Beginning of the adaptation activities for the validation bench BV. 

BV permits to validate the flight software; It simulates all the Data handling Unit (DHU) interfaces in opened or 

closed loop. The simulation in closed loop is possible for the AOCS chain. The flight software for validation is 

loaded from a computer to the DHU. 

BV permits also to validate all the satellite functional chains. It simulates all command/control interfaces of 

these chains in opened or closed loop. The simulation in closed loop is possible for the AOCS, thermal and 

electrical chains and the payload. 

BV activities requires a payload functional interface model (see section 6.1.2.1). 

• Beginning of the software modifications and the system data base update. 

• Beginning of the activities for the launch vehicle adapter realization. 

Moreover, the first Launch Coupled Analysis is performed. 

The necessary inputs are : 

• A new payload Data Package (DP 2) containing at least : 

• A Payload Description Document (synthesis document) 

• A Payload Interface Control Document 

• A set of mathematical models 

• A CAD model 

• Two Finite Element Models (one physical model and one reduced model) 


• A Payload Design, Development and Validation Plan 

• A Payload Verification Plan 

• A Payload AIV requirements 

The Critical Design Review concludes the phase C; that permits to freeze the design and to start realization. 




10.1.4 PROTEUS PHASE D 

During this phase, ALCATEL SPACE manages the modifications defined in Phase C. In this phase, the platform 

equipment units and the long lead items are supplied . Then, ALCATEL SPACE leads the platform Assembly 

Integration and Tests. And as soon as the payload flight model is received and accepted, ALCATEL SPACEleads the 

satellite Assembly Integration and Tests. 

Notice : The payload shall be entirely qualified before ALCATEL SPACE delivery. At satellite level, only payload health 

check and acceptance tests are planned. 

ALCATEL SPACE delivers to the Customer: 

• the satellite numerical models for launch vehicle, 

• the system data base, 

• the satellite flight model on launch site, 

• a « Telemetry Tracking and Command suitcase » is available; it permits to check the compatibility between 

satellite and ground segment, 

A satellite simulator is delivered to the ground segment in order to validate the operational interfaces (optional). 

The Authority in charge of the payload delivers to ALCATEL SPACE: 

• a payload functional interface model for the validation bench (BV) adaptation, this model shall be fully 

representative in term of electrical and functional interfaces 

• a payload flight model, 

• the payload ground support equipment is available, during all satellite activities : from payload delivery to 

launch. 

10.1.5 PROTEUS PHASE E 

During this phase, ALCATEL SPACE supports CNES for the launch campaign and the flight acceptance. 

In PROTEUS standard, CNES operates the satellite in Toulouse. 

The Customer operates the mission center. 




10.2 SATELLITE DOCUMENTATION 

The documentation relating to the PROTEUS mission can be divided in the main following parts : 

• the applicable documents (inputs) and interfaces documents 

• the management documents 

• the product assurance documents 

• the documents describing the development and validation logic 

• the system description and performances documents 

• the justification documents 

• the documentation relating to operations 

• the documentation dealing with the mission centre/PROTEUS Generic Ground Segment interfaces 

The purpose of this chapter is to list the main documents delivered for each activity quoted above. In the following 

tables, « Supplier » designates ALCATEL SPACE or CNES. 

The publication dates are satellite events dates. 

10.2.1 APPLICABLE DOCUMENTS AND INTERFACES DOCUMENTS 

10.2.1.1 Satellite and system 

Title Issued by Publication 

date 

Mission System Requirements Customer 

T0 Pre-phase B 

T0 phase B 

Satellite Specification 

Customer 

or 

CNES 

Satellite to Ground Interfaces CNES 


T0 phase B 

T0 phase C 

T0 phase B 

T0 phase C 

System Test Requirements Customer CDR 

System Test Plan Customer CDR 

for information 

Comments / Purpose 

this document refers to the standard 

“PROTEUS Satellite to Ground interfaces” 

10.2.1.2 Launch vehicle interfaces 

The launch vehicle choice and the launch configuration are definitive at the beginning of phase B (BDR). The main 

documents exchanged with the launch vehicle authority are listed in the User’s Manual of the chosen launch vehicle. 




10.2.1.3 Payload interfaces 

PL - 10.2.1 -1 

The payload Supplier shall delivered all the documents mentioned in this Table as issued by the Customer 


date 

PROTEUS User’s Manual (PUM) Supplier Standard 

Payload Interfaces Requirements 

Description (PID) 

Payload Design and Interfaces 

Specifications (PDIS) 

Payload Interfaces Specifications 

Compliance and Verification Matrix 

Customer T0 Pre-phase B 

Supplier 

Customer 

Payload CAD models (electronic files) Customer 

STA mechanical mathematical 

models (electronic files) 

Payload mechanical mathematical 

models (electronic files) 

Supplier 

Customer 

Payload budgets Customer 

BDR 

PDR 

CDR 

CDR and with 

PL FM delivery 


T0 phase B 

T0 phase C 

Standard 

PDR 


T0 phase B 

T0 phase C 

after correlation 


T0 phase B 

T0 phase C 

Payload description document Customer T0 phase B 

T0 phase C 

Payload Interface Control Document 

(PICD) 

Customer 

T0 phase B 

T0 phase C 

and at every 

change 

Some of these documents are more precisely described in the following paragraphs. 


it gives the interfaces & environment 

specifications for a payload based on 

PROTEUS before the first PDIS issue. 

this document describes the payload 

interfaces 

this document is written from the PUM and 

PID and is the applicable interface 

document for the payload 

this document allows to check the PL 

interfaces compliance 

These inputs are necessary to lead satellite 

accommodation under the fairing, field of 

view verification ... These models shall be 

provided with the associated drawings. 

these inputs are necessary to lead payload 

mechanical analyses. These models will be 

provided with a description document. 

these inputs are necessary to lead satellite 

mechanical analyses. These models shall 

be provided with a description document. 

Mass properties, power (for each lines and 

each PL mode), data rates 

This documents provides a description of 

the payload and gives at least : 

• Payload and mission overview 

• Description of each architecture 

(mechanical, thermal, electrical, 

• 

command and control) 

And other specific features 

PICD is composed of at least 

• Payload Interfaces Data sheet 

• Grounding scheme 

• Interfaces Descriptions Drawings 

• ... (see section 4.1.1) 




10.2.1.3.1 PROTEUS User’s Manual (PUM) 

PROTEUS User’s Manual is a generic document describing more particularly interfaces specifications between 

PROTEUS Satellite and Payload. This document is a standard document and is available at the beginning of project. 

It is written for three PROTEUS User groups : System Prime, Mission Center Prime and Payload Prime. 

10.2.1.3.2 Payload Design and Interfaces Specification (PDIS) 

PDIS is a specifying document, particular for each mission. It is a PROTEUS User’s Manual specific adaptation and its 

table of contents is the same as the PUM one. PUM is a reference document for PDIS. Adaptations to the generic 

specifications, if needed for the mission, are described in PDIS. 

It deals with interfaces constraints imposed by the Satellite Prime to the Payload Prime. Specifications result from 

Platform, satellite and system levels. 

This document is updated in each design phase thanks to payload data packages and satellite level analyses. 

10.2.1.3.3 Payload Interface Control Document (PICD) 

Payload Prime describes payload interfaces in PICD. PICD is the answer to the PDIS, that is to say that for each PDIS 

issue, there is a PICD issue. 




10.2.2 MANAGEMENT DOCUMENTS 


date 

Management plan Supplier PDR 

Review reports Supplier PDR 

CDR 

FAR 

Schedule report Supplier PDR, each 6 

weeks 

Payload Deliverable Items list Supplier PDR, CDR 

10.2.3 PRODUCT ASSURANCE DOCUMENTS 


At each review, a report is published 


date 


Satellite Product Assurance Plan Supplier PDR It gives the rules applicable to the satellite 

follow up 

Satellite Configuration Items Data 

List 

Supplier PDR 

CDR 

FAR 

It gives the references of the documents 

useful for the supply, the fabrication, the 

tests and deliveries of products 

Satellite Qualification Status list Supplier CDR, FAR It gives the qualification state of each 

equipment, sub-system, system. 

Satellite Deviations List Supplier PDR, CDR, FAR It lists the deviations emitted by the 

suppliers. 

Payload Material and Mechanical 

Part List 

Customer CDR - 2 months 

Payload Process List Customer CDR - 2 months 

Payload EEE part List Customer PDR - 2 months 

OBSW Software Quality Assurance 

Plan (to be discussed) 

Supplier PDR It gives the rules to implement the software 

Payload Reliability analysis Customer PDR, CDR 

Payload Safety analysis Customer PDR, CDR it studies the design conformity regards to 

the rules applicable on launch sites 

Satellite Material and Mechanical 

Part List 

Supplier CDR 

Satellite Process List Supplier CDR 

Satellite EEE part List Supplier PDR 

Satellite Reliability analysis Supplier PDR, CDR 

Satellite Safety analysis Supplier PDR, CDR 




10.2.4 DOCUMENTS RELATING TO DEVELOPMENT AND VALIDATION LOGIC 

10.2.4.1 Development and Validation plans 


date 

Payload Design, Development & 

Validation Plan 

Payload Verification Plan & Test 

Matrix 

Satellite Development & Validation 

plan 

10.2.4.1.1 Verification Plan 

Customer T0 phase B 

T0 phase C 

Customer T0 phase C 

T0 phase D 

Supplier PDR 

CDR 


This document defines the development, 

qualification model, tests philosophy to 

comply with satellite interfaces, 

environment and planning and also how to 

validate it. 

See here below. 

This documents describes the philosophy 

chosen for satellite validation (cf. PUM & 

PDIS) 

The Verification Plan shall describe how the Payload will verify each requirement of the PDIS and PUM (Test, 

Analysis...). 

10.2.4.1.2 Test Matrix 

In addition to the Verification Plan, a Test Matrix shall be prepared that summarizes all the tests that will be 

performed on the payload. The purpose of the matrix is to provide a quick reference to the contents of the test 

program in order to prevent the deletion of a portion thereof without an alternate means of accomplishing the 

objectives; it has the additional purpose of ensuring that all flight hardware has seen environmental exposures that 

are sufficient to demonstrate acceptable workmanship. In addition, the matrix shall provide a review of the 

qualification heritage of hardware. All flight hardware, spares and prototypes (when appropriate) shall be included 

in the matrix. 

The Test Matrix shall be prepared in conjunction with the initial Verification Plan and shall be updated as changes 

occur. 

10.2.4.2 Specifications of facilities for satellite integration and system tests 


date 


Technical Requirements of payload Supplier PDR This documents lists the requirements for 

simulator 

payload simulator 

Payload functional and interfaces Customer CDR Numerical models useful for validation 

model for validation benches 

bench 

Payload Simulator User’s Manual Customer for BV 

adaptations 




10.2.5 TEST PLANS, ASSEMBLY INTEGRATION AND TESTS 


date 

Payload AIT Plan Customer PDR 

CDR 

Payload AIV requirements Customer T0 phase B 

T0 phase C 

Payload User’s Manual Customer T0 phase C 

and at PL FM 

delivery 

Payload End Item Data Package Customer CDR and with 

PL FM delivery 

Satellite AIT plan Supplier PDR 

CDR 

Payload integration (on satellite) 

procedures 

Supplier for satellite AIT - 

3 months 


Integration and test plan at payload level 

Defines Payload verification and tests at 

satellite level 

These inputs are useful for satellite AIT 

Data package to be delivered with the 

payload flight model 

Integration and test plan at satellite level 

For Customer approval to verify Payload 

safety during satellite AIT. 

Satellite End Item Data Package Supplier FAR Data package to be delivered with the 

satellite flight model. 

10.2.6 SYSTEM DESCRIPTION AND PERFORMANCES 


date 

Satellite executive summary Supplier PDR 

CDR 

Satellite budgets and margins Supplier PDR 

CDR 

FAR 

In Orbit Test Plan Supplier PDR 

CDR 

Satellite mechanical ICD Supplier PDR 

CDR 

Satellite electrical ICD Supplier CDR 

Functional synoptic Supplier PDR, CDR + 

every major 

modification 


Presentation of SL architecture 

It presents mass, fuel, power & energy, link, 

TM & TC, pointing & stability, reliability & 

availability budgets and margins at satellite 

level. 




10.2.7 JUSTIFICATION DOCUMENTS 


date 

Design Verification Matrix and Customer PDR 

conformity to PDIS requirements 

CDR 

PDR 

Satellite justification files Supplier CDR 

FAR 

Design Verification Matrix and 

conformity to satellite requirements 

Design Verification Matrix and 

conformity to On board/Ground 

requirements 

10.2.8 SYSTEM DATA BASE AND OPERATIONS 

Supplier PDR 

CDR 

FAR 

Supplier PDR 

CDR 

FAR 


date 

Command control and Operations Supplier CDR, 

User Manual 

FAR - 6 months 

Satellite Telemetry plan Supplier PDR, CDR, FAR 

Satellite Telecommand Plan Supplier PDR, CDR, FAR 

Satellite Simulator User’s Manual Supplier phase D 


It gives all the specific justification 

documents for each functional chains. The 

generic ones are given in the PROTEUS 

justification documents. 


10.2.9 MISSION CENTRE/PROTEUS GENERIC GROUND SEGMENT INTERFACES DOCUMENTATION 


date 


Proteus User’s Manual Supplier NA 2 chapters deal with the generic ground 

segment and its interfaces 

Command Control Ground Segment 

Description 

Supplier NA 

Ground Segment System Description Customer CDR 

PGGS adaptation specification Supplier CDR 

PGGS Interfaces Supplier CDR ∆ with respect to the PUM 

Network interfaces architecture Supplier CDR 




10.3 PROTEUS PROVISION 

10.3.1 PAYLOAD MECHANICAL MATHEMATICAL MODELS 

PL - 10.3.1 -1 

These models shall be delivered by Payload Supplier to lead mechanical and thermal analysis at the 

following dates: 

• Beginning of the Satellite pré phase B 

• Beginning of the Satellite phase B (T0) 

• Beginning of the Satellite phase C (T0 + 4) 

• After payload environment tests (correlated model) 

The requirements are given in section 4.6 

10.3.2 PAYLOAD CAD MODELS 

PL - 10.3.2 -1 

These models shall be delivered by Payload Supplier at the following dates: 

• Beginning of the Satellite pré phase B 

• Beginning of the Satellite phase B (T0) 

• Beginning of the Satellite phase C (T0 + 4) 

The requirements are given in section 4.6 

10.3.3 PAYLOAD FUNCTIONAL INTERFACES MODEL (EM OR PAYLOAD SIMULATOR ?) 

PL - 10.3.3 -1 

These interfaces shall be delivered by Payload Supplier for validation bench (BV) adaptation phase (T0+11). 

A preliminary definition of this model is given in section 6.1.2. 

10.3.4 DELETED 

10.3.5 STAR TRACKER PROVISION 

A Star Trackers Assembly composed of a STA flight model equipped with mechanical breadboard of 2 STRs, with its 

associated STA User’s Manual, will be provided by ALCATEL SPACE to Payload Supplier for the Payload environment 

tests. A STA User’s Manual model shown in appendix D permit to see in particular the STA integration procedure 

requirement. 

The Star Trackers Assembly (Flight model) will be integrated on Satellite at ALCATEL SPACE. 

The Star Trackers wiring harness (between STR and platform) will be provided by ALCATEL SPACE. 

PL - 10.3.5 -1 

The Star Tracker Assembly thermal control harness (monitoring and active thermal control) shall be provided 

by the Payload Supplier with ALCATEL SPACE requirements. 

10.3.6 PLATFORM/PAYLOAD INTERFACES WIRING PROVISION 

10.3.6.1 Platform/Payload power interfaces 

Bracket H01 (platform/payload power interfaces bracket) and wiring from platform to this bracket are provided by 

ALCATEL SPACE. 




PL - 10.3.6 -1 a 


Wiring from this bracket to the Payload shall be made by Payload Supplier, but connectors (P01 to P03 and 

P06 to P08), on payload side, will be provided by ALCATEL SPACE. 

10.3.6.2 Nominal & Redundant Platform/Payload TM/TC interfaces 

Wiring from platform to the Platform/Payload TM/TC interfaces brackets H02 and H03 will be provided by ALCATEL 

SPACE. 

Brackets H02 and H03 will be also provided by ALCATEL SPACE. 

PL - 10.3.6 -2 a 

Wiring from these brackets to the payload shall be provided by Payload Supplier, but connectors (J01 to J07 

and J09 to J12), on payload side, will be provided by ALCATEL SPACE. 

10.3.7 THERMAL MLI 

PL - 10.3.7 -1 

MLI protection for H02 and H03 brackets shall be provided by the Payload Supplier and it shall be possible 

to dismount these MLI several times. 

MLI protection between the payload and the platform will be provided by ALCATEL SPACE. 

PL - 10.3.7 -2 

The Payload shall contain attachment points for the previous MLI between the payload and platform. 

PL - 10.3.7 -3 

Deleted 

10.3.8 PLATFORM/PAYLOAD INTERFACES SCREWS 

PL - 10.3.8 -1 

For each type of delivered screws, the associated torquing tools shall be provided by the Payload Supplier 

10.3.8.1 Payload interface screws 

PL - 10.3.8 -2 

The 4 x 4 platform/payload interface M8 screws shall be provided by the Payload Supplier. 

10.3.8.2 STA interface screws 

PL - 10.3.8 -3 

The 8 STA/payload interface M5 screws shall be provided by the Payload Supplier. 

10.3.8.3 Electrical brackets interface screws 

PL - 10.3.8 -4 

The 6 bracket/payload interface M5 screws shall be provided by the Payload Supplier for each bracket (H02 

and H03). 







PRO.LB.0.NT.003.ASC Issue. 06 rev. 03 Page: C.1 

APPENDIX – C 

STANDARD STA IDS 



Title PUM 6.3 Appendix C – STA IDS 

Title CALIPSO Star Trackers Assembly Reference PRO-LBP-O-IC-3060-ASP 

Issue 1 Issue Date 18/02/2003 

Revision 0 Revision Date 

Authors Christophe DUPLAY 

Product code 

Issue / Revision Change notice summary / Applicability 

1 / 0 ORIGINE 

Use ALT-RETURN for add a line in a same cell. 

Mechanical architect : 

Technical Manager : 

Thermal architect : 

Procurement manager : 

Electrical architect : 

Command/Control architect : 

Configuration manager : Quality manager : Payload manager : 

Page C.2 29/11/04

Title Reference Issue 

SED16 Interface Control Document for 

Proteus 

Reference list PUM 6.3 Appendix C – STA IDS 

Issue 

date 

Revision 

Rev 

Date 

PRO.SOD.IS.SED1600010 11/06/02 A 11/06/02 

Description 

Page C.3 29/11/04

PL_Mechanics PUM 6.3 Appendix C – STA IDS 

MECHANICAL CHARACTERISTICS 

Envelope DIMENSIONS in mm: C.G LOCATION in mm: MASS in kg 

L 474,00 +/- 5,00 CGx: 0,00 +/- 5,00 Nominal Mass 11,400 

W: 466,00 +/- 5,00 CGy: 4,00 +/- 5,00 Mass Variation 

DIA: +/- CGz: -179,00 +/- 5,00 Mass Dispersion 

H: 377,00 +/- 5,00 Maximum Mass 12,500 

Allocated Mass 

Nominal inertia provided in STA reference frame axes 

INERTIA in m^2.kg 

Ixx: 0,6 +/- 0,05 Ixy: 0 +/- 

Iyy: 0,6 +/- 0,05 Ixz: 0 +/- 

Izz: 0,2 +/- 0,05 Iyz: 0 +/- 

-6 

MATERIAL OF HOUSING AND SURFACE FINISH: Housing material : aluminum honeycomb with carbon face sheets (CTE < 2.10 m/m/°) 

NUMBER OF CONTACT POINTS 8 Contact points material: PERMAGLASS ME 730 

CONTACT AREA OF EACH POINT in cm^2: 0,265% of the baseplate area: 

FLATNESS OF CONTACT AREA in mm: 0,10 

ROUGHNESS OF CONTACT AREA in microns rms: 3,20 

EIGENFREQUENCY in Hz > 150 Hz TIGHTENING THICKNESS in mm: 19 (see annexed sketch) 

Page C.4 29/11/04

PL_Therm PUM 6.3 Appendix C – STA IDS 

THER MAL CHAR ACTER ISTICS 

For radiative part of the thermal s izing, the following datas s hall be cons idered 

Cf ICD Drawing: drawings with all the dimens ions define every S TA coating. 

The following table completes the drawings 

Thermal-optical features Temperatures limits (°C) 

Coating Coating type eir amin amax op. mode non-op. mode 

area (BOL) (EOL) Tmin Tmax Tmin Tmax 

MLI 0,77 0,32 0,49 adiabatic equilibrium with environment 

R adiative Area (S S M) 0,76 0,10 0,16 -15,00 30,00 -40,00 30,00 

For conductive part, the following datas shall be considered 

Global thermal conductive coupling 0.04 W/°C (0.005 W/°C per contact points) 

Thermal-optical features Temperatures limits (°C) 

Type eir amin amax op. mode non-op. mode 

(BOL) (EOL) Tmin Tmax Tmin Tmax 

S TA s tructure NA NA NA -15,00 30,00 -40,00 30,00 

Page C.5 29/11/04

Drawings PUM 6.3 Appendix C – STA IDS 

Page C.6 29/11/04


Page C.7 29/11/04


END OF APPENDIX 

Page C.8 29/11/04

PRO.LB.0.NT.003.ASC Issue. 06 rev. 03 Page: D.1 

APPENDIX – D 

STA USER’S MANUAL 

(this model correspond to a STA flight model equipped with mechanical breadboard of 2 STRs 





1. SCOPE 1 

2. APPLICABLE DOCUMENTATION 1 

2.1 PROJECT SPECIFIC DOCUMENTATION 1 

2.2 GENERAL DOCUMENTATION 1 

2.3 ACRONYMS 1 

3. STA MASS MODEL 1 

3.1 STA General description 1 

3.2 Mass model representativity 1 

4. STA MASS MODEL INSTRUMENTATION 1 

5. STA MASS MODEL INTEGRATION PROCEDURE 1 

5.1 STA mass model integration on Payload 1 

5.2 Tightening procedure with thermal washers: 1 

5.3 Electrical connection & Harness routing 1 

5.4 STA Grounding on PL 1 




1. SCOPE 

The present document describes the Star Tracker Support Assembly (STA) mass model and its integration procedure. 

2. APPLICABLE DOCUMENTATION 

2.1 PROJECT SPECIFIC DOCUMENTATION 

NA 

2.2 GENERAL DOCUMENTATION 

NA 

2.3 ACRONYMS 

CTA: Active Thermal Control 

STB : Requirement specification 

N/A : not applicable 

Nida : Honeycomb 

STR : Star Tracker 

STA : Star Tracker Assembly 

TML : Total Mass Loss 

CVCM : Collected Volatile Condensable Material 

PL : Payload 

3. STA MASS MODEL 

3.1 STA GENERAL DESCRIPTION 

STA is composed of 2 STR mass model and the STA flight carbon structure. 

The structure is composed of the primary structure, the structure grounding, the thermal control connector mounting on its 

bracket (H20). 




TOTAL MASS : TBD maximum calculated mass 

3.2 MASS MODEL REPRESENTATIVITY 

The STA mass model is structurally flight representative with STR mass models. 

• Mass 

• COG and Inertia 

• First modal frequency and first structural mode: 142 Hz, lateral oscillation. 

• Geometrical interface 

• Fixation component (insulating washers) 

• Electrical connectors (with savers for STR connector and screw lock). 




4. STA MASS MODEL INSTRUMENTATION 

ALCATEL needs 5 accelerometers located as shown on the figures hereafter: 

STR2:Triaxe 

STR3:Triaxe 

STR4:Triaxe 




STR1:Triaxe 

STR5:Triaxe 

Nota : For commodity reasons, the STR’s shown on the figure are the flight CAD representation. 

Sensor Type Location Orientation 

ST1 3 axes compatible with Sine On STA structure 

Parallel to STA Frame 

and acoustic Test frequency As close as possible from 

range.. 

STR1 attachment point 

ST2 3 axes compatible with Sine On STA structure 


and acoustic Test frequency As close as possible from 

range.. 

STR2 attachment point 

ST3 3 axes compatible with Sine On STA baseplate 


and acoustic Test frequency As close as possible vertical 

range. 

panel. 



and acoustic Test frequency On +X STA axis close to the 

range. 

baseplate border 



and acoustic Test frequency On - X STA axis close to the 

range. 

baseplate border 




5. STA MASS MODEL INTEGRATION PROCEDURE 

5.1 STA MASS MODEL INTEGRATION ON PAYLOAD 

See Annexe figure 1 

The M5 titanium screws are provided by Payload supplier. The mini tension required is defined in PL-3.4.6-1. 

The tightenig torque TBD is given by Payload supplier 

The thermal washers (16 units+ 16 spare) are provided, by Alcate l(rep STA01). 

The aluminium washer (10 units) are provided, by Alcatel (rep 344). 

The onduflex washer (10 units) are provided, by Alcatel (rep 324). 

5.2 TIGHTENING PROCEDURE WITH THERMAL WASHERS: 

First the torque is applied to each screw. 

After 30 minutes the torque shall be applied a second time. 

After 48 hours the torque shall applied a third time. 

5.3 ELECTRICAL CONNECTION & HARNESS ROUTING 

See annexe figure 2 : 

Savers are accommodated on STR connectors in order to not damage the STR cable. 

The material of STR female connectors screw-locks is gilded brass. 

The material of H20 female connectors screw-locks is inox female connectors screw-locks. 

The tightening torque of connector screw-lock is : 0.33 N.m. 

5.4 STA GROUNDING ON PL 

The ground braids are mounted on the payload. The PL supplier shall connect the 2 ground braids with the 2 dedicated stud as 

shown in next figure. 

(TBC) 

(TBC) 

(TBC) 

(TBC) 




ANNEX 





Drawings PUM 6.3 Appendix D – STA User’s Manual 

Figure 1 : STA Integration (CALIPSO Exemple)



Drawings PUM 6.3 Appendix D – STA User’s Manual 

Figure 2 : STA Electrical connexion and cable routing (CALIPSO Exemple)


END OF APPENDIX

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