DESIGN AND ANALYSIS OF THE THERMAL CONTROL SYSTEM FOR SPACE TECHNOLOGY 5

1. DESIGN AND ANALYSIS OF THE THERMAL CONTROL SYSTEM FOR SPACE TECHNOLOGY 5 David Neuberger Swales Aerospace Incorporated Beltsville, Maryland 15th Annual Thermal and Fluids Analysis Workshop Pasadena, California August 30th – September 3rd, 2004

2. ST-5 Mission Concept Space Technology 5 is NASA’s pathfinder for highly capable, low-cost small spacecraft, miniaturized subsystems, and constellation mission operations. Research-Quality Spacecraft “The ST5 project shall demonstrate the ability to achieve accurate, research-quality scientific measurements utilizing a nanosatellite with a mass less than 25 kg. ” Micro-Satellite Design and Build “The ST5 Project shall design, develop, integrate, test and operate three {one} full service spacecraft, each with a mass less than 25kg, through the use of breakthrough technologies. ” Constellation Mission “The ST5 project shall execute the design, development, test and operation of multiple spacecraft to act as a single constellation rather than as individual elements. ”

3. ST-5 Mission Summary • Full Functional Autonomous Spacecraft with Integrated Technology • Science Grade Magnetic Sensitivity (~ 1 nT) • Mass: ……..…. 25Kg • Size: ………….. Diameter ~ 53 cm (Solar panel peak-to-peak) Height ~48 cm (Antenna tip to antenna tip) • Power: ……….. ~20-25W at 9-10V • ~7-9 Ah Battery • Uplink: ………. @ 1Kbps / Downlink: @1Kbps or 100Kbps (X-Band) • Data Storage: .. 20 Mbyte • Spin Stabilized at Separation (~25 RPM After Deployments) • Deployments: Magnetometer • Radiation Tolerant: 100 Krad-Si TID SPACECRAFT • Launch Vehicle: Pegasus XL • Launch Location: Vandenberg AFB, Lompoc, CA • Orbital Injection: Sun-Synchronous Polar Elliptical Orbit (300km x 4500km) • Three Spacecraft carried on Spacecraft Support Structure as Prime Payload LAUNCH • 3-Spacecraft Constellation • 3-Month Design Life • 136 min Orbit Period • 10-30 Minutes Ground Contact Three Times Per Day • Autonomous Constellation Management / “Lights Out” Operations MISSION

4. Baseline Orbital Elements Launch Timeframe:          February-March 2006 Launch Site:                Vandenberg AFB, Lompoc, CA Mission Duration:             90 days Eclipses:                             None due to earth shadow, March 29 eclipses on 2 - 3 orbits due to moon shadow Perigee:                            300 km Apogee:                           4500 km Inclination:                       105.6 deg (sun synchronous) Period:                                  136 min Number of orbits/day:  about 10.5 RAAN:                            42 deg or so for Feb 15 launch, increasing 1 deg/day for launch later in launch window (full sun 6 AM - 6 PM) Launch Argument of Perigee:   160 deg Rotation of Apsides: -1.2 deg/day (apogee rotates towards South Pole)

5. 25 Kg Research Spacecraft X-Band Antenna Variable Emittance Surface (radiator and electronics) Miniature Spinning Sun Sensor Nutation Damper Low Voltage Power Subsystem (Li-Ion battery, triple junction solar cells) X-Band Transponder Deployment Boom Cold Gas Micro-Thruster CULPRiT Chip (on C&DH card) MiniatureMagnetometer (sensor and electronics)

6. ST-5 Spacecraft • Developed by GSFC • Description • Built within tight volume and mass constraints • Low-power and low voltage • ~53 cm x 48 cm • Integral card cage structure (for C&DH, PSE) • Key performance parameters • Mass less than 25 kg • Spacecraft-induced magnetic field effects as measured at the magnetometer sensor location less than 10 nT (d.c.), 5 nT (a.c.) • Ground Testing • Component testing per ST5-495-007 (including magnetics) • FLATSAT for electrical integration of engineering models • Spacecraft-level functional and environmental testing (including magnetics) • Flight testing • Operations of the spacecraft during the 90-day mission (including magnetometer measurement of s/c magnetic field) • Future applicability: s/c useful as-is or re-use components

7. Spacecraft Layout (Deployed) 50.8cm Mag Sensor NutationDamper Mag Electronics VEC Controller #2 C&DH VEC Controller #1 3.0 cm PSE 73.1 cm Mag Boom X-Band Antenna Battery +Zsc +Xsc VEC Radiator #1 TCE Transponder Electronics 10.5 cm HPA X-BandAntenna PressureTransducer 27.0 cm Thruster 28.6 cm Diplexer Sun Sensor PropellantTank +Ysc +Xsc 10.5 cm VEC Radiator #2 Thruster{Nozzle Exit} +Zsc

8. Spacecraft Thermal Design (1 of 3) • Sizing Conditions • The thermal design was established with the assumed condition that the spacecraft is spinning with the spin axis of the spacecraft normal to the ecliptic plane ±5°. • The thermal design was sized assuming a spacecraft internal heat dissipation of ~20 watts for hot case and ~13 watts for cold case. • Electrical power is subtracted off solar arrays. • An electrically conductive coating with a high emittance is used for the radiators, e.g. NS43C. • Black paint/anodize is used to provide a high emittance interior. • Heat in and out dominated by solar energy absorbed by and radiated from body-mounted solar array panels; robust design. • Highest possible apogee assumed for cold case, lowest possible apogee assumed for hot case. • The ST-5 spacecraft TCS will utilize passive thermal control techniques (coatings, thermal conductors and isolators, insulation blankets, etc)

9. Spacecraft Thermal Design (2 of 3) • Most components are mounted to the interior of the top and bottom decks • VEC ECUs and the TCE are mounted to the side of the card cage • Nutation Damper is mounted to sidewall of spacecraft • Multi-Layer Insulation • The gaps between adjacent solar array panels and the panels and the spacecraft will be closed out using multi-layer insulation. • A 2 layer Kapton “skirt” will be used around the base of the X-Band Antenna. • The top and bottom decks will be covered with MLI on the external surfaces with a window cutout for radiators • The Magnetometer Sensor Head will be wrapped with MLI. • The rigid boom segments and root adapter will be wrapped with MLI. • The battery will be wrapped with MLI on five sides and a low e film on the sixth side (facing the deck).

10. Spacecraft Thermal Design (3 of 3) • To minimize sources of heat loss/gain, some components will be conductively isolated. • VEC Radiators and the X-Band antenna will be isolated from the spacecraft using G10 standoffs • The eight solar array panels will be conductively isolated from spacecraft using low conductivity mounting bracket, but radiatively coupled to spacecraft using a high emittance coating on sidewall (substrate already has a high emittance) • The Magnetometer boom will be conductively isolated from the spacecraft and the sensor. • The battery has been designed to be conductively isolated from the spacecraft. • Other methods are used to increase conduction: • Nusil and Teflon will be used for the High Power Amplifier and Transponder • PSE and C&DH cards are heat sunk to the cardcage using a wedgelok along the right and left edges. • Relatively large bolts are used to provide good contact conductance at bolted interface between the cage and the decks.

11. Thermal Analyses Modeling Philosophy • Orbit parameters and operational scenarios were provided by the Project. • 7 on orbit cases • 7 launch cases • How Design Margin is Implemented • Use conventional conservative hot and cold case modeling assumptions to overestimate the predicted temperature extremes. • Qualify components 10°C beyond their design limits. • Incorporate additional design margin by keeping predicts at least 5°C from design limits, which results in at least 15°C margin on predicts. • Due to swings in temperatures, multiple runs are completed in Sinda until a quasi-steady state is achieved. • Temperature predicts represent max and min temperatures for various cases.

12. General Assumptions • Sun synchronous polar orbit • Spacecraft spins with spin rate 25 rpm • Spin axis perpendicular to sun 5° • FMH = 18 W/m2, applied for 10 minutes before or after perigee *Since Nusil was used for these boxes, 300 Btu/hr/ft2 was used for the transponder and HPA.

13. Material Properties

14. Optical Properties • Values were received from Coatings Committee • Values for other materials were derived from the composite of several materials/coatings or came from testing/analysis and can be provided upon request.

15. Cases * 3 times per day

16. Power Assumptions • Solar Array Power draw: • The sum all component powers plus diodes is subtracted off solar array. • Only valid as long as power is supplied by solar array only and/or battery is not charging. Power pulled off of array is limited to 24.4 Watts. • Diodes on entire time. • Essential bus on entire time.

17. Cold Case Steady State Power

18. Hot Case Steady State Power

19. Model (1 of 2)

20. Model (2 of 2)

21. Predicts

22. Predicts

23. Predicts

24. Predicts: Hot Short Transmit Case

25. Predicts: Hot Short Transmit Case

26. Predicts: Cold Operation Case

27. Predicts: Cold Case Operation

28. Predicts • For the hot case, the radiator, solar arrays, and most internal components run at ~37°C. • For the cold case, most internal components run at ~0°C. • All components have at least 5°C margin on their hot and cold operational and survival limits except the following: • MEMS power profile needs to be updated. • HPA’s orbital average is ~32°C, however it spikes up to 47°C when transmitting. • Had to open up the radiator to get positive margin on the HPA in the hot case, but was limited by TCE in cold case.

29. Small Model Hints • ST5 much smaller than most SC • Same laws of thermodynamics • Blanket effective emittance higher than standard SC • Using 0.03 to 0.1 instead of standard 0.005 to 0.03 for larger SC. • Will find out at spacecraft thermal balance test. • Errors • A few square inches to a 100 in2 radiator is a much higher % than it is for a 1000 in2 one. • Low energy balance. Tracking down tenths of Watts instead of Watts. Not as critical as a cryo cooler where one is worried about milliwatts. • Pay close attention to details. Some components (ex. connectors) may be similar size to that in a large SC. They would represent a larger % of area on a small SC and can’t be neglected. • Check sensitivity to critical parameters • Sensitivity to power for ST5 is 0.5°C per Watt. • You do this with all spacecraft designs anyway but may be critical with the smaller ones.