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

David NeubergerSwales Aerospace Incorporated

Beltsville, Maryland

15th Annual Thermal and Fluids Analysis WorkshopPasadena, California

August 30th – September 3rd, 2004

ST-5 Mission Concept

“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. ”

“The ST5 project shall demonstrate the ability to

achieve accurate, research-quality scientific

measurements utilizing a nanosatellite with a mass

less than 25 kg. ”

“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. ”

Micro-Satellite Design and Build

Research-Quality Spacecraft

Constellation Mission

Space Technology 5 is NASA’s pathfinder for highly capable, low-cost small spacecraft, miniaturized

subsystems, and constellation mission operations.

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

• 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

• 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” OperationsM

ISSION

LAUNCH

SPACECRAFT

Baseline Orbital Elements

Launch Timeframe:          February-March 2006Launch Site:                Vandenberg AFB, Lompoc, CAMission Duration:             90 daysEclipses:                             None due to earth shadow, March 29

eclipses on 2 - 3 orbits due to moon shadowPerigee:                            300 kmApogee:                           4500 kmInclination:                       105.6 deg (sun synchronous)Period:                                  136 minNumber of orbits/day:  about 10.5RAAN:                            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 degRotation of Apsides: -1.2 deg/day (apogee rotates towards South

Pole)

25 Kg Research Spacecraft

MiniatureMagnetometer

(sensor and electronics)

DeploymentBoom

Variable Emittance Surface (radiator and electronics)

X-Band Transponder

Miniature Spinning

Sun Sensor

CULPRiT Chip(on C&DH card)

Low Voltage Power

Subsystem (Li-Ion battery,

triple junction solar cells)

Cold Gas Micro-Thruster

Nutation Damper

X-Band Antenna

ST-5 Spacecraft• Developed by GSFC

• Description

o Built within tight volume and mass constraintso Low-power and low voltageo ~53 cm x 48 cmo Integral card cage structure (for C&DH, PSE)

• Key performance parameters

o Mass less than 25 kgo Spacecraft-induced magnetic field effects as measured at the magnetometer sensor

location less than 10 nT (d.c.), 5 nT (a.c.)

• Ground Testing

o Component testing per ST5-495-007 (including magnetics)o FLATSAT for electrical integration of engineering modelso Spacecraft-level functional and environmental testing (including magnetics)

• Flight testing

o 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

Spacecraft Layout (Deployed)

NutationDamperMag Electronics

C&DH

Battery

PropellantTank

PSE

VEC Controller #1

Sun Sensor

VEC Controller #2

Transponder Electronics

HPA

VEC Radiator #2

X-Band Antenna

73.1 cm

3.0 cm

50.8cm

Mag Boom

Mag Sensor

+Zsc+Xsc

+Ysc+Xsc

+Zsc

VEC Radiator #1

28.6 cm

Thruster{Nozzle Exit}

TCE

PressureTransducer

DiplexerThruster

X-BandAntenna

10.5 cm

27.0 cm

10.5 cm

Spacecraft Thermal Design (1 of 3)• Sizing Conditions

o 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°.

o The thermal design was sized assuming a spacecraft internal heat dissipation of ~20 watts for hot case and ~13 watts for cold case.

o Electrical power is subtracted off solar arrays.

o An electrically conductive coating with a high emittance is used for the radiators, e.g. NS43C.

o Black paint/anodize is used to provide a high emittance interior.

o Heat in and out dominated by solar energy absorbed by and radiated from body-mounted solar array panels; robust design.

o 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)

Spacecraft Thermal Design (2 of 3)• Most components are mounted to the interior of the top and

bottom deckso VEC ECUs and the TCE are mounted to the side of the card cage

o Nutation Damper is mounted to sidewall of spacecraft

• Multi-Layer Insulationo The gaps between adjacent solar array panels and the panels and the

spacecraft will be closed out using multi-layer insulation.

o A 2 layer Kapton “skirt” will be used around the base of the X-Band Antenna.

o The top and bottom decks will be covered with MLI on the external surfaces with a window cutout for radiators

o The Magnetometer Sensor Head will be wrapped with MLI.

o The rigid boom segments and root adapter will be wrapped with MLI.

o The battery will be wrapped with MLI on five sides and a low e film on the sixth side (facing the deck).

Spacecraft Thermal Design (3 of 3)• To minimize sources of heat loss/gain, some components will be

conductively isolated.

o VEC Radiators and the X-Band antenna will be isolated from the spacecraft using G10 standoffs

o 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)

o The Magnetometer boom will be conductively isolated from the spacecraft and the sensor.

o The battery has been designed to be conductively isolated from the spacecraft.

• Other methods are used to increase conduction:

o Nusil and Teflon will be used for the High Power Amplifier and Transponder

o PSE and C&DH cards are heat sunk to the cardcage using a wedgelok along the right and left edges.

o Relatively large bolts are used to provide good contact conductance at bolted interface between the cage and the decks.

Thermal Analyses Modeling Philosophy

• Orbit parameters and operational scenarios were provided by the Project.

o 7 on orbit cases

o 7 launch cases

• How Design Margin is Implemented

o Use conventional conservative hot and cold case modeling assumptions to overestimate the predicted temperature extremes.

o Qualify components 10°C beyond their design limits.

o 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.

o Temperature predicts represent max and min temperatures for various cases.

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.

Parameter Hot Case Cold Case

Solar Constant (W/m2) 1418 1285Albedo 0.35 0.25

IR (W/m2) 265 208Perigee 300 km 300 kmApogee 1500 km 4000 kmShadow (%) 0 0Period (hr) 1.7 2.2

19.1 (Nominal) 13.2 (Nominal)37 (During Transmit) 16.7 (VECs on)

MLI e* 0.03 0.1Optical Properties EOL BOLSun Angle (to spin axis) 5° 0°Conduction at Interface per type of screw Min MaxConduction at S/C and box Interface (Btu/hr/ft2) 50 300Conduction at S/C and Comm. Box I/F (Btu/hr/ft2) 300* 300*

Total Power (Watts)

Material Properties

Material k(W/cm°C) Cp(J/Kg°C) LocationAluminum 6061 1.8 962 Decks, boxes, heat sink, etc.Aluminum A356 1.6 962 Card Cage AssemblyAluminum 7075 1.2 962 Magnetometer TangM55J 0.277 922 BoomM46J 0.201 922 Solar Array PanelsG-10 0.003 ** Antenna and VEC standoffsEccofoam SH-2 0.0002 ** Qual Helix AntennaUltem 7801 0.024 900 Snubbers, boom bracketCopper 3.9 418 Circuit boards, wiresNitrogen (Cv) N/A 733 PropellantTitanium 6Al-4V 0.073 648 Boom partsBeryllium Copper 1.3 418 Boom hingesStainless Steel 0.163 ** Bolts

** Arithmatic node or conduction path

BOL BOL EOL EOL

A276 on Raydome White 0.28 0.87 0.4 0.86Buffed Aluminum Yellow 0.19 0.06 0.24 0.04Irridite Yellow 0.18 0.18 0.5 0.05GBK Grey 0.49 0.81 0.51 0.78Black Kapton Black 0.93 0.89 0.95 0.85Solar Cells Blue 0.89 0.8 0.93 0.77Beryllium Copper (Oxidized) Turquoise 0.82 0.1 0.89 0.08Black Anodize Black 0.7 0.82 0.86 0.78Clear Anodize White (Black in pics) 0.33 0.8 0.45 0.72Copper Orange 0.32 0.05 0.55 0.03Gold Yellow 0.18 0.04 0.2 0.03Kapton Orange 0.41 0.8 0.55 0.76Z306 on Raydome Red 0.96 0.91 0.98 0.88NS43G Red 0.2 0.91 0.46 0.88Stainless Steel Turquoise 0.42 0.16 0.47 0.11Ultem 7801 Black 0.86 0.83 0.91 0.81VDA tape Purple 0.1 0.06 0.18 0.03Z306 Pink 0.94 0.91 0.96 0.88Z307 Pink 0.94 0.88 0.97 0.84Kapton on Solar Array 0.86 0.85 0.9 0.83TefRam (Measured on Ti) Turquoise 0.73 0.27 0.77 0.24

Coating Color

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.

Cases

* 3 times per day

CaseApogee Altitude

Sun Angle (Relative to Spin Axis)

MLI ε*Un Regulated Bus Voltage

Powered On or Off

VEC MagEssential

BusTransmitter

Cold Safehold 4000km -25° 0.1 6.0 V Off Off On No

Hot Safehold 1500 km +25° 0.03 8.4 V Off Off On No

Cold Survival 4000km 0° 0.1 6.0 V Off Off On No

Cold Operational 4000km 0° 0.1 8.4 V On 1, Off 4 On On No

Hot Operational (short transmit)

1500 km + or – 5°0.03

8.4 V On On On 18 min*

Hot Operational (long transmit)

4000 km + or – 5°0.03

8.4 V On On On 30 min*

Power Assumptions

• Solar Array Power draw:

o The sum all component powers plus diodes is subtracted off solar array.

o 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.

Cold Case Steady State Power

Hot Case Steady State Power

Model (1 of 2)

Model (2 of 2)

Predicts

-60

-40

-20

0

20

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On-Orbit Predicts

Op Limits

Non-Op Limits

<5°C Margin on Predicts

-60

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On-Orbit Predicts

Op Limits

Non-Op Limits

<5°C Margin on Predicts

Predicts

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-60

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On-Orbit Predicts

Op Limits

Non-Op Limits

<5°C Margin on Predicts

Predicts

-120

-100

-80

-60

-40

-20

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20

40

60

80

100

120

140

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Column 9Column 8

On-Orbit Predicts

Op Limits

Non-Op Limits

Post-Deployment Limits

<5°C Margin on Predicts

Black Kapton hotter than GBKRadiators clearly visible

Poorly coupled irridite

Black Kapton hotter than GBKRadiators clearly visible

Poorly coupled irridite

Predicts: Hot Short Transmit Case

Predicts: Hot Short Transmit Case

Black Kapton hotter than GBK

Radiators clearly visible Black Kapton hotter than GBK

Radiators clearly visible

Predicts: Cold Operation Case

Predicts: Cold Case Operation

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:

o MEMS power profile needs to be updated.

o 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.

Component Case Margin

HPA Hot Short Transmit 2.8°C

MEMS Radiator Cold Operation 1.4°C

TCE Cold Operation 3.4°C

Small Model Hints

• ST5 much smaller than most SC

o Same laws of thermodynamics

o 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.

o 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.

o 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.