Sounding Rocket Technology Demonstrator for Small Satellite ...

Sounding Rocket
Technology Demonstrator
for Small Satellite Launch
Vehicle Project
Purdue University
John Tsohas, Lloyd J. Droppers,
Stephen D. Heister
1
April 26, 2006

Purpose & Goals
• Develop technologies critical to launch vehicle
development
• Phase 1: Design/Build Sounding Rocket
Technology Demonstrator
– 5 lb payload to 25,000 ft
– Test-bed for critical sub-systems including
ground support, propulsion, structures,
separation, guidance, navigation and control.
– Sequential flight validation of sub-systems.
• Phase 2: Design of small satellite launch vehicle
– 10 lb payload to LEO (93 mi. orbit).
– Use technologies demonstrated in Phase 1.
– 3 stages, hybrid propulsion (H 2 O 2 /HTPB), low
propellant mass fraction design
2

Overview of Sounding Rocket Demonstrator
3
• Propulsion:
– 250 lbf thrust 90% H 2
O 2
/HTPB hybrid,
pressure-fed system.
– H 2
O 2
: high density Isp, non-cryogenic,
non-toxic, safe propellant handling,
reduces complexity/operations costs.
• Structure:
– Carbon-fiber composite airframe
– Aluminum oxidizer pressure vessel
– Fiber-glass internal structure,
honeycomb fins.
• Recovery:
– Dual pyrotechnic parachute
deployment (primary/drogue)
– Redundant recovery system avionics.
• Ground Support Equipment:
– Remote fill/draining of hydrogen
peroxide to and from vehicle
– Remote control and monitoring of
sounding rocket launch operations.
• Guidance and Control:
– Liquid injection thrust vector control
(LITVC) technology in development
– Flight verification in subsequent flights

Propulsion
4
• Sub-scale, 25 lbf Thrust Engine
– Hot fire testing to characterize
HTPB regression rate with
90% H 2 O 2
– Validation of internal ballistics
engine design code
– Data used for design of full
scale engine
• Full-scale,170 lbf Thrust Engine
– Flight weight engine design
– 440 psia MEOP, 5.25 avg. O/F
ratio.
– Internal phenolic liner for
casting HTPB and insulation
– Butt-seal design, and
secondary o-ring seal
– Silica-phenolic nozzle
– Carbon-filled EPDM and RTV
insulation
– Consumable Catalyst Bed
used for ignition of H 2 O 2 /HTPB
– 218 second ISP (sea-level,
moderately over-expanded)
Chamber Pressure
Thrust
Oxidizer Mass Flow Rate
Initial O/F ratio
Predicted average c*
Fuel Grain Outer Port
Diameter
Fuel Grain Length
440
175
0.68
5.25
4900
3.5
10
[psi]
[lbf]
[lbm/s]
[ft/s]
[in]
[in]
Injector Assembly
Combustion Chamber
Top Cap
Consumable
Catalyst Bed
Aluminum
Motor Casing
HTPB Fuel
Grain
Aft Mixing
Section
Paper Phenolic
liner
Silica Phenolic
Nozzle
Viton O-ring
Nozzle
Retaining Ring

5
Sub-scale 25 lbf Thrust Engine Hot Fire Testing

Structure Sub-system
6
• Carbon-fiber, 6” diameter
airframe
• Aluminum oxidizer
pressure vessel
(MEOP=600 psia)
• Burst pressure F.S.
2 X MEOP
• Proof tested to
1.5 X MEOP
• Vibration tests scheduled
• Fiber-glass nose cone
and internal structure
design
• Honeycomb fin design
• Performed Finite Element
Analysis on major
components

Ground Support Equipment
• Remote fill/drain of H 2 O 2 to and from vehicle
• Remote control of solenoid valves with Labview software
• Regulated nitrogen pressure used to transfer H 2 O 2 from
storage tank to rocket oxidizer tank through series of
solenoid valves (600 psia MEOP).
• Rocket oxidizer tank temperature and pressure monitored
to ensure safe condition of H 2 O 2
• Fail-safe design: In event of power outage, valves return to
normal positions, automatically venting tanks and draining
oxidizer from launch vehicle
7

Design Philosophy of Small Launch Vehicle
• Primary driver for universities is to provide student
body with education and practical engineering
experience to design, build and fly small launch
vehicle into LEO
• ‘In series’ manufacturing and flight verification of
each stage individually will help work out kinks in
sub-systems, without placing entire integrated
vehicle at risk.
• Serial development will allow each generation
(class) of student designers to work from inception
to completion of project while still being students
(each launch vehicle stage).
• Relatively low propellant mass fraction will allow for
student design, and local manufacturing.
8

Conceptual Design of Small Launch Vehicle
• Trajectory, aerodynamics,
and vehicle sizing codes
developed for initial design
• Design constraints:
– 10 lb payload, to
minimum 93 mi LEO
orbit.
– Pressure fed system
design
– 10,000 lbf max. thrust
based on Purdue test
capabilities.
– ISP and engine
perfrormance data from
Phase 1 testing.
9
M tot stg1 M prop stg1 M i stg1 M tot stg2 M prop stg2 M i stg2 M tot stg3 M prop stg3 M i stg3 M payload M GLOW
5397 4156 1241 786.9 605.9 180.7 81.2 62.5 18.7 10.0 6275.1
86%
GLOW
77%
STG1
23%
STG1
12.5%
GLOW
77%
STG2
23%
STG2
1.3%
GLOW
77%
STG3
23%
STG3
0.16%
GLOW
(mass in lbs)

Conceptual Design of Small Launch Vehicle
10
• Stage 3:
– SRM propulsion
– Satellite release mechanism
– Avionics module
• Stage 2:
– 1,230 lbf thrust, 98%
H 2 O 2 /HTPB hybrid motor
– Burn duration: 157 sec
– ISP: 320 sec(avg)
– Chamber pressure: 200 psia
– LITVC thrust vector control
– Carbon-fiber composite
structure
• Stage 1:
– 8,790 lbf thrust, 98%
H 2 O 2 /HTPB hybrid motor
– Burn duration: 123 sec
– ISP: 260 sec
– Chamber pressure: 600 psia
– LITVC thrust vector control
– composite structure.
– Parachute recovery
– Novel altitude compensation
Stage 2 mass breakdown:
Ullage Mass, lbm 15.1 8.4%
Oxidizer Tank mass, lbm 16.2 9%
Engine Mass, lbm 28.1 15.6%
Pressurant Mass, lbm 5.5 3%
Pressurant tank, lbm 16.8 9.3%
Structural Mass, lbm 18.1 10%
Miscellaneous mass, lbm 27.1 15%
Mass Growth, lbm 53.8 29.7%
Stage 2 Total Mass, lbm 180.7 100%
Stage 1 mass breakdown:
Ullage Mass, lbm 103.9 8.37%
Oxidizer Tank mass, lbm 201.6 16.2%
Engine Mass, lbm 281.1 22.6%
Pressurant Mass, lbm 69.7 5.6%
Pressurant tank, lbm 230.7 18.6%
Structural Mass, lbm 124.1 10%
Miscellaneous mass, lbm 124.1 10%
Mass Growth, lbm 105.9 8.5%
Stage 3Total Mass, lbm 1241 100%

Conclusions
11
Phase 1:
• Development of sounding rocket technology demonstrator
for flight verification of critical sub-systems for small
launch vehicle design (Phase 1)
• Sounding rocket currently in assembly phase.
• Launch from NASA Wallops Flight facility
• Addition of guidance and control sub-systems on follow-on
flights.
Phase 2:
• Small Satellite Launch Vehicle in conceptual design
phase.
• Design requirement based on capabilities and needs of
university.
• Design philosophy based on ‘serial’ development program.
Incremental flight validation of each individual stage.
• Code development and trade studies in progress.