NANOSAT THERMAL MANAGEMENT WORKSHOP

Sept 2003

Charlotte GerhartMechanical Engineer

Charlotte.Gerhart@kirtland.af.mil

505-846-2438

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Spacecraft Thermal Management

• Objective: To describe what, why, when, and how spacecraft thermal management is accomplished.

– What is Thermal Management?

– Why is Thermal Management Important for Spacecraft?

• Temperature Effects

• Orbital Environment

– How are Thermal Systems Designed?

• Quick Check Calculations

• Numerical Modeling

• Accepted Orbital Thermal Modeling Standards

• Examples

– Thermal Design/Model Inputs

– Other Thermal Design Considerations

– AFRL’s Role

– Thermal Engineers Toolbox

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What is Thermal Management?

– Assuring that a system operates within it’s temperature range

• Automobiles

• Computers

• Refrigerator

• Aircraft Fuselage

• Alaska Oil Pipeline

• Spacecraft

– OR determine temperature limits and make sure the system can handle the induced loads

• Sidewalks

• Railroads

• Power Lines

• Bridges

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Why is Thermal Management Important for Spacecraft?

• Prevalent Misconceptions:– Space is cold, getting rid of heat is no problem!– It runs just fine in the lab– If it gets too hot just put a fan on it– Just put a heater on it if it’s cold

• Reality:– Heat transfer and rejection is via radiation in space, which is very

different than terrestrial systems which are convection dominated– The sun is VERY hot and produces a lot of heat that can be absorbed

(1350W/m2)– The earth is a “cool” (~250K) body that strongly effects earth orbiting

missions AND reflects some solar energy back out to space– There is no air in space for a fan to blow around unless you are

willing to build a pressure vessel (very heavy, prone to leaks), and you still have to eventually dump the heat by radiation.

– Heaters require power, usually mostly during eclipse = bigger battery = more mass/volume

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Temperature Effects

• Failure– Expansion of deployable panel to stick in frame– Freezing of lubricant

• Device out of tolerance/calibration• Thermal fatigue

– Cracking, breaking solder joints• Thermal distortions

– Stresses, bending

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The Earth Orbital EnvironmentE

CL

IPS

E

Waste Heat

Incident SolarReflected Solar

Earth IR

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Thermal System Design and Analysis

• Iterative process with increasing fidelity– Steady state thermal balance– Transient analysis with minimal number of nodes and surfaces– More detailed analyses with better information, focus on critical

components• Analysis code MUST include radiation and conduction as well as orbit

propagation– view factors– solar inputs– earth IR (earthshine) – reflected (albedo)

• NASA standard: – Sinda (finite difference solver) – Trasys (view factors)

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Steady State Estimations

• Quick check average environmental sink temperatures for earth, 3-axis stabilized cube (Aerospace Corporation)

Beta 0 Beta 90 Beta 0 Beta 23.5Nadir Face 246K 250K 204K 206KZenith Face 211K 0K 210K 206KVelocity Face 221K 180K 210K 206KAnti-Velocity 221K 180K 210K 206KNorth Face 176K 292K 0K 223KSouth Face 173K 179K 0K 0K

Cube SurfaceLEO (500nm = 925km) GEO (19360nm = 35835km)

Beta 0

Beta 90

Assumptions: Surface properties abs = 0.2, emis = 0.8. 1400W/m2 solar, 250W IR, 0.33 albedo orbit average fluxes

Beta is the angle between the earth’s orbit plane(normal to the page in this example) and that of the satellite

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Quick Check Example

• 500W radiated from 1m2 earth facing panel in LEO, beta=0, emissivity=0.8 (assume back side is perfectly insulated, abs=0 = neglecting solar and earth inputs)

• General radiative heat transfer equation:

Q = A(Ts4 - T

4) [1]Where: = Stefan-Boltzmann Constant = 5.67e-8 W/m2*K4

= emissivityA = area, m2

Ts = Surface Temperature, K

T = Sink Temperature, KQ = Heat Input, W

Re-arranging Equation 1 and solving:

Ts = (Q/(A* * ) + T )1/4 = 324K = 51ºC• Can be used to get first order estimates of temperatures, radiator and

heater sizes

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Thermal System Design and Analysis

• Other first order checks:– System thermal capacity for transient response time ROM

• 10kg Aluminum (875 J/kg*K) + 2kg Copper (380 J/kg*K) = 9150 J/K

9150 J/K = 9150 W*s/K = 152 W for 1 minute for 1K temperature change

• Energy balance on inputs, internal generation, and output for rate of change

• Next step is building FEM’s to do steady state and/or transient analysis– Lumped nodes, surface properties, contact/mounting

conductances– Trade different properties to obtain desired temperatures

• Surface coatings, insulation, mounting, heaters, etc.• AFRL uses: SINDA and RADCAD

– Have I-Deas for Structural, not Thermal

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Example Analysis (UN1)

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Example Analysis Results (UN1)

NadirMax (C)

NadirMin (C)

Anti-NadirMax (C)

Anti NadirMin (C)

MSDS 22 -38 40 -34MSDSBattery

-4 -4 -2 -2

TopEmerald

5 -12 0 14

BottomEmerald

0 -12 0 -14

Orion -2 -18 -2 -18

MSDS with Black Paint

NadirMax (C)

NadirMin (C)

Anti-NadirMax (C)

Anti NadirMin (C)

MSDS -7 -35 -14 -44MSDSBattery

-13 -13 -22 -22

TopEmerald

17 -2 -12 -22

BottomEmerald

-9 -20 -17 -27

Orion -13 -26 -21 -32

MSDS with White Paint

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Other Thermal Considerations

• Temperature gradients– Plasma thrusters (hot) next to IR sensor (cold)

• Surface degradation– Space environment– Handling– Outgassing– Thrusters

• Cost and handling procedures– Abrade surfaces, toxic materials, ground processing

• Thermal shorts– Harnessing, compressed MLI

• Need thermal balance testing to verify model – Requires vacuum - preferably thermal vacuum + solar sources -

and lots of temperature sensors (ground only)

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Thermal Toolbox

• Coatings– Black Paint (emis=.95, abs=.85) -GEO, internal faces– White Paint (emis=.9, abs=.2) -LEO– Silver Teflon (emis=.9, abs=.1)

• Insulation– MLI: properties vary with # of layers and specific application– Beta cloth: shuttle/ISS material (touch temperatures)

• Interface conduction– G-10, ceramic paper: low conductivity interface material– Calgraph (carbon paper), thermal grease, metallic foils, etc.: for

high conductivity gap filler• Heat transfer

– High conductivity metals - aluminum, copper– An-isotropic composites - graphite's, carbon fibers– Heat pipes, loop heat pipes, capillary pumped loops– Pumped fluid loops or circulators

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Thermal Toolbox

• Thermal energy storage– Thermal mass– Phase change materials (paraffin's for room temperature)

• Heat rejection– Louvers– Variable emissivity coatings– Heat pumps

• Cryogenic systems– Thermal electric coolers, cryocoolers, cryogenic radiators

• Suggested component temperature ranges:– Operational: -10C to +60C– Non Operational: -40C to +100C– Deployment, heat sensitive actuation: -50C to +150C

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AFRL’s Role

• Responsible for system in shuttle through post ejection deployment• Have vested interest in the success of University missions• Provide thermal models to NASA for Shuttle manifesting

– Flight rules– Limited thermal control (heaters, thermostats)

• Need detailed information to build STS paylaod level models• Can provide satellite level analysis and review/oversight as desired

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Summary

• Thermal design and analysis essential for spacecraft– Keep all spacecraft components in their temperature ranges during

all mission phases– Ambient operation is generally not indicative of space

• Iterative process throughout design and mission• Analysis codes must solve conductive and radiative heat transfer, and

as be able to simulate the orbital environment• Models are only as good as the information used • No simple “one size fits” all solutions• AFRL is here to provide pre-deployment survival as well as help as

needed to assure mission success• Suggested component temperature ranges:

– Operational: -10C to +60C– Non Operational: -40C to +100C– Deployment, heat sensitive actuation: -50C to +150C