Avionics Preliminary Design Review - University of Maryland

Avionics 

ENAE 788D Design Project 


Establishing a Recurring Human 

Presence on the Moon 

Preliminary Design Review 

University of Maryland, Fall 2012 

Block, Henninger, Rotunda

Overview 

• Preliminary Design Review of avionics, 

software and simulation for a low-cost lunar 

lander 

– Link Budget and Communications 

– Sensors 

– Simulation (Design/Build/Test/Evaluate) 





Link Budgets 

• Link budget analysis conducted for the following 

cases: 

– Ku band direct to earth 

– S band direct to earth 

– Ka band to relay satellite at Lagrange 2 

– Ku band from relay satellite to Earth 

– UHF Omni-directional to EVA suits 

• Link margins >7 dB desired to ensure 99.5% link 

availability 





Assumptions 

• Maximum operational distances: 

– Direct to earth = 384,000 km 

– To Lagrange 2 = 60,000 km 

– L2 to earth = 444,000 km 

– Spacecraft to EVA Suits = 1 km 

• Data Rates: 

– Ka band = 100 Mbps 

– Ku band = 25 Mbps 

– S band = 100 kbps 

– UHF = 50 kbps 

• Ground station antenna diameter = 15m 

• Relay satellite specs from Boeing TDRS-J 

• All other assumptions from standard spreadsheet 





Ku-Band Direct to Earth 





S-Band Direct to Earth 





Ka-Band to Lagrange 2 Relay Satellite 





Ku-Band L2 Relay Satellite to Earth 





UHF Omni-directional to Relay Suits 





Link Budget Conclusions 

• Baseline design is a 0.3 meter diameter 

transmitting antenna on spacecraft 

• Ku-band direct to earth requires a 1 m dia. 

transmitting antenna, a 50 m dia. receiving 

antenna or a reduced data transfer rate of 2.25 

Mbps; reduced rate is the only practical option 

• S-band direct to earth meets all requirements 

under nominal conditions, but at a lower data 

transfer rate (100 kbps) and a higher power 

requirement than Ku-Band (40 W vs. 20 W) 





Link Budget Conclusions 

• Ka-band relayed through L2 satellite to earth in 

Ku-band meets all requirements in nominal 

condition assuming a relay satellite comparable 

to Boeing TDRS-J, includes cost of additional 

satellite 

• UHF omni-directional to EVA suits meets gain 

requirements with 1 cm dia. transmission and 

receiving antennas for EVA’s up to 1 km; ample 

room to scale for missions with long distance 

EVA’s 





Proprioceptive Sensors 

• Pressure Transducers 

• Control Position Indicators 

– Attitude Control 

– Altimeters 

– Gyroscope 

– Strain Gauges 

– Magnetometer 

• Concern for South Atlantic Anomaly 

• Fluid Flow Sensors 

– Internal Cabin Gas monitoring 

– Propulsion Systems 

• Thermal Sensors 

– Internal Cabin 

– Propulsion systems 

• Space Crew Sensors 

– Radiation Sensors (Dosimeters) 

– Gas Leak Sensors 

• Each tied to alarm system 



Adapted from [5] 



Exterioceptive Sensors 

• Position Indicators 

– Star Trackers 

– Telemetry Devices 

– Landing Gear Position Indicators 

• Laser Altimeters (LIDAR) 

– Cameras 

• Analyzing Devices 

– Tied to probability and calculation software 

• Detecting Devices 

– Micrometeoroid Sensors 

– Ablation Sensors 

• Camera 

– Horizon Sensors 

– Infrared Sensors 





Examples of Sensors 

• Pressure Sensors 

– Internal cockpit sensor 

• Critical for maintaining atmospheric conditions 

• Pressure Transducer provides extremely low weight[2] 

• LIDAR support for landing struts 

– Redundancy for comparative analysis 

– Serious weight issue (100 kg per system) [6] 

• Cameras- Everywhere 

– Infrared 

– Pan Tilt Zoom 

– Additional masking or filter features 

• Star tracker 

– Readily available COTS brand 

• Radiation Sensor 

– Simple badge or dynamic sensor 



SpaceMicro’s μStar 



Processing Power 

• Proton400k Or RAD 750 

• Requires suitable Error Detection and 

Correction (EDAC) 

• Multiple computer banks for redundancy 



Adapted from [2] 



Sensors 

• Radiation Tested Equipment a major concern 

– 3 main scenarios of testing [4] 



• Total Ionizing Dose: Co 60 gamma or Co-3 MeV electrons 

( Linac or VdG ) 

• Displacement Damage: Protons (10 20 MeV ), Neutrons 

(1 MeV ), Electrons (3 5 MeV 10-MeV), MeV), 3- 

• Single Event: Heavy Ion Accelerator (ESA Louvain HIF), 

Proton Accelerator (ESA PSI PIF) ESA- ESA-Cf 252 CASE 

laboratory system. Cf-‘CASE’ 1-VdG) MeV) 



Proprioceptive Sensor List 

Sensors Type Criticality Frequency of sampling Weight (kg) Amount Total 

Pressure Transducers Moderate 9.1 Hz 2 5 10 

Position Indicators 25 1 25 

Attitude Control High - - 

Altimeters Slight 60 Hz - - 

Gyroscope Moderate 60 Hz - - 

Strain Gauges Moderate 60 Hz - - 

Magnetometer Moderate 333 mHz - - 

Fluid Flow Sensors 

Internal Cabin Gas Monitors High 900 Hz 1 4 4 

Propulsion Systems High 900 Hz 1 6 6 

Thermal Sensors 

Internal Cabin Slight 900 Hz 0.25 4 1 

Propulsion Systems High 900 Hz 0.25 6 1.5 

Space Crew Sensors 

Radiation Sensors (Dosimeters) High N/A 1 6 6 

Gas Leak Sensors High 900 Hz 0.25 6 1.5 

Total Weight 55 





Exterioceptive Sensors List 

Sensors Type Criticality Frequency of sampling Weight (kg) Amount Total 

Position Indicators 

Star Trackers Moderate 60 Hz 1.85 2 3.7 

Telemetry Devices High 60 Hz 4 2 8 

Landing Gear Position Indicators 

LIDAR High 100 Hz 100 2 200 

Cameras Moderate 60 Hz 0.5 20 10 

Analyzing Devices 

General Purpose Computers High 3 GHz 10 5 50 

Total Weight 271.7 



Total weight of all sensors 326.7 kg 



Design/Build/Test/Evaluate Projects 

1. Cabin Interior 

– Living and working environment during transit and 

on lunar surface 

– Object and workstation placement 

2. Lunar landing visuals and operations 

– Landing gear must be visible from cabin window 

3. Ingress/Egress 

– Difficulty level of ingress/egress on lunar surface 

while wearing spacesuits 





Cabin Interior – Overview 

• Limited interior space creates close-quarters 

living and working environment 

• Interior includes: 

– 3 launch seats, food storage, waste management 

system, 3 spacesuits, 3 crew members, additional 

stowage 

• Sleeping and working arrangements must 

provide comfort level expected of 10-13 day 

mission 





Cabin Interior – Research Objectives 

• Size and placement of each interior item 

– Examine various possible combinations of item placement 

– Simulate living and working. Determine which placements work well and 

which are hazardous or inconvenient 

• Working environment 

– 3 crew members must work as a team. Simulate working together 

during each mission phase 

– Each mission phase may have unique requirements, i.e. launch seats for 

take-off, window visibility for lunar landing 

– Ensure workstations are accessible during each phase, promote crew 

cohesion and contribute to mission success 

• Living and sleeping arrangements 

– Transit and lunar phases offer different challenges due to gravity 

– Floor space must be sufficient for 3 sleeping crew members 

– Determine comfort level of occupying launch seats for sleeping 





Cabin Interior – CONOPS 

• Construct shell of cabin interior size 

– Material is not important, i.e. can be cardboard 

– Cut holes for windows and hatch (include door) 

• Create mockup of each interior item 

– Find/create seats, food, simulated spacesuits 

– Most important is to size each item as accurately as possible 

– Items must be easily movable to allow for multiple 

placement possibilities 

• Simulate all phases and aspects of mission 

– Launch, landing, sleeping, eating, working together 

– Evaluate each for feasibility and optimal environment 





Cabin Interior – Mockup 

Interior Height: 2.37 m 

Landing 

Controls / 


Food 

Storage 



Interior Diameter: 3.13 m 

Waste 

Management 

Collapsible 

Seats (3x) 



Lunar Landing – Research Objectives 

• Lunar landing is an especially critical mission 

phase which depends on crew for success 

• All crew members must be gainfully employed 

– Ensure window, workstation, avionics placement allows 

each crew member to contribute optimally to the 

success of this mission phase 

• Windows must allow for 360° lateral visibility 

• Sight lines such that landing gear feet are visible 

during landing 

– Enable crew members to have maximum chance of 

avoiding landing hazards 





Lunar Landing – CONOPS 

• Elevate cabin interior mockup such that landing 

gear can be placed beneath 

– Place cabin interior mockup on a stand to accurately 

simulate the height of lunar landing module and takeoff 

propulsion module 

– Landing gear mockup must be accurate length and 

angle to evaluate sight lines from window 

• Landing hazards 

– Simulate various placements of landing hazards such as 

holes and rocks 

– Determine crew’s chances of hazard detection and 

avoidance based on placement 





Lunar Landing – Mockup 



Sight Lines (3 windows 

evenly spaced around SC) 

41.9° 



Ingress/Egress – Research Objectives 

• Extra Vehicular Activity (EVA) is a primary purpose 

of overall mission 

• Crew members must be able to don spacesuits, 

egress and ingress easily and readily 

• Investigate how each of the following affects this 

mission phase: 

– Hatch size and placement 

– Ladder deployment and stowage 

– Spacesuits and cabin interior 

• Ensure that crew members are not overly 

encumbered during this phase and are able to 

focus on mission 





Ingress/Egress – CONOPS 

• Donning spacesuits inside cabin 

– All crew members must be able to don suits prior to cabin 

depressurization 

– Mockup spacesuits must simulate actual suit size and mobility 

level to ensure crew can don them properly inside cabin 

• Hatch size and placement 

– Hatch must be appropriately sized and positioned above cabin 

floor to allow feet-first egress 

• Ladder deployment 

– Mockup ladder must be of material to support weight of one 

crew member plus suit 

– Mockup ladder must be compressible such that it can be stored 

inside cabin and deployed easily outside hatch while wearing suit 





Ingress/Egress – Mockup 

Hatch Diameter: 1.0 m 



EVA Hatch 



Mass Budget – Crew Vehicle 



Component Mass (kg) 

Crew Systems 1323 

Power, Propulsion and Thermal 1342 

Structure 682 

Sensors 327 

Total 3674 

• 1121 kg remaining for additional mass 



References 

• [1] http://lhcb-elec.web.cern.ch/lhcb-elec/html/radiation_hardness.htm 

• [2]http://trs-new.jpl.nasa.gov/dspace/bitstream/2014/15130/1/00-1141.pdf 

• [3] http://www.spacemicro.com/pdfs/Proton400k-8X_Datasheet_v3.2.pdf 

• [4] http://esamultimedia.esa.int/docs/industry/SME/2003/Space-Component/ESA- 

Training-Radiation-ESTEC_May03.pdf 

• [5] http://www.lanl.gov/orgs/mpa/mpa11/Panel-Hunter-NASA.pdf 

• [6] Radiation Testing of Semiconductor devices for Space Electronics 

http://web.mst.edu/~umrr/cf043.pdf 

• [7] http://ww.gim-international.com/files/productsurvey_v_pdfdocument_11.pdf 

• [8] Koon, W. S. Low Energy Transfer to the Moon. Celestial Mechanics and Dynamical 

Astronomy. September 2001, Volume 81, Issue 1-2, pp 63-73. 

• [9]http://www.eumetsat.int/Home/Main/Satellites/GroundNetwork/GroundStations/in 

dex.htm 

• [10] http://www.boeing.com/defensespace/space/bss/factsheets/601/tdrs_hij/tdrs_hij.html 

• [11] http://tdrs.gsfc.nasa.gov/assets/files/PressKits/TDRS%20J.pdf 

• [12] Maral, Gérard, and Michel Bousquet. Satellite Communications Systems: Systems, 

Techniques, and Technology. 5th ed. Chichester: Wiley, 2011. Print.

Avionics Preliminary Design Review - University of Maryland

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