Thermal Control Systems
Thermal Control Systems
Temperature and heat are intimately related since
temperature is a measure of heat energy
• Three common temperature scales
• Kelvin – most useful since zero energy equals zero
temperature
– freezing water = 273.15 K (boiling water 373.15 K)
• Celsius – 0o Celsius is freezing point of water, 100oC
is boiling point of water, -273.15oC is absolute zero
• Fahrenheit – +32oF is freezing point of water, 212oF is
boiling point of water, -459.67°F is absolute zero
Thermal Control Systems
A quick review of two of the basic laws of thermodynamics
are important in the basic review of thermal systems
• Zeroth law of thermodynamics relates two systems being
in equilibrium with a third implies they are also in
equilibrium with each other for isolated (closed) systems
– This means that heat always flows from hot to cold
• First law of thermodynamics equates change in heat of an
isolates system is equivalent to the same change in
energy and temperature
– This means heat-energy is conserved in an isolated
system, and that the change in heat-energy can be
measured by temperature, and that different forms of
energy and heat are equivalent
Thermal Control Systems
• Second law of thermodynamics characterizes two
systems in contact that tend towards equilibrium
temperature and energy
– Verifies that heat always flows from hot to cold
– Identifies why perpetual-motion machines are not
possible
Thermal Control Systems
Heat sources and sinks
• Heat can be supplied and removed by a wide variety
of processes and mechanisms
• Heat can be transferred in space by five physical
mechanisms
– Conduction
– Radiation
– Convection (limited)
– Evaporation
– Condensation
Thermal Control Systems
Heat transfer
– Conduction - the transfer of heat energy by
physical contact between a solid, liquid, or gas
– Radiation - the transfer of heat energy by
electromagnetic radiation through emission or
absorption
– Convection - the transfer of heat energy by mass
transfer of liquid or gas (requires gravitational field
or acceleration)
– Evaporation – cooling by the evaporation of a
liquid or solid (solid to gas is called sublimation)
– Condensation – heating by the condensation of
gas to liquid or to solid (deposition), or liquid to
solid (fusion)
Thermal Control Systems
Electromagnetic radiation heating and cooling
• Both heating and cooling occur with electromagnetic
(EM) radiation depending on the equivalent
temperature of the radiation and the temperature of
the physical surface
• Absorbed heat energy – heat gain efficiency is
measured by the absorptivity (α) of the material
surface at specific EM frequencies
• Emitted heat energy – heat loss efficiency is
measured by the emissivity (ε) of the material surface
at specific EM frequencies
Thermal Control Systems
Absorptivity α
Emissivity ε
Thermal Control Systems
Kirchkoff's law - a good emitter is a good absorber at a
particular frequency, or for a specific range of
frequencies
• Equivalent to a material having both high emissivity
and absorptivity (or both low) within a specific band
• Generally true for only a few materials
Thermal Control Systems
Blackbody radiation - any object emits heat energy
which is characteristic of the peak temperature of the
material
• The temperature-frequency curve for a blackbody has
the same shape for all blackbody objects, with the
peak of the curve representing the peak temperature
• A cooler blackbody curve is represented by a smaller
area and is shifted to lower energy than a warmer
body, but retains the same characteristic shape
Thermal Control Systems
Blackbody curve
for various
temperatures
including the Sun
(yellow) and the
Earth (red)
Radiant emittance
is power radiated
per area, and often
called intensity
(radiated)
Areas under the
curves represent
total radiated
power
Thermal Control Systems
Passive systems - no active devices (requiring energy)
are used for heating, cooling or transferring heat
• Examples are radiator fins, insulation, coatings
Active systems - mechanical and/or electrical systems
used for heating, cooling, or heat transfer
• Examples are cooling loops, heaters, and
refrigerators
Thermal Control Systems
The primary source of energy in the solar system is
electromagnetic radiation from the Sun. Although the
gravitational potential energy of the Sun is significant, it
cannot be used for conventional heating, cooling, and
electrical power for spacecraft.
Planet
Distance (AU)
Solar radiation (W/m²)
Perihelion
Aphelion
maximum
minimum
Mercury
0.3075
0.4667
14,446
6,272
Venus
0.7184
0.7282
2,647
2,576
Earth
0.9833
1.017
1,413
1,321
Mars
1.382
1.666
715
492
Jupiter
4.950
5.458
55.8
45.9
Saturn
9.048
10.12
16.7
13.4
Uranus
18.38
20.08
4.04
3.39
Neptune
29.77
30.44
1.54
1.47
Thermal Control Systems
• Approximate temperature range in the solar system
– Average solar energy available at 1 AU is 1340 Watts/m2
– Composed of 50% IR, 40% Vis, 10% UV
Distance from Sun
Solar Exposure (wavelength
and material dependent)
In Shadow (wavelength and
material dependent)
Near Mercury (0.3
AU)
800 K (980oF)
50 K (-280oF)
Near Venus (0.7 AU)
500 K (440oF)
50 K (-280oF)
Near Earth (1.0 AU)
400 K (260oF)
50 K (-280oF)
Near Mars (1.5 AU)
300 K (80oF)
20 K (-280oF)
Near Jupiter (5.3 AU) 100 K (-280oF)
Near Pluto (40 AU)
20 K (-420oF)
20 K (-420oF)
20 K (-420oF)
Thermal Control Systems
• Temperatures on or in a spacecraft are influenced by
the sum of heat inputs (sources) and heat outputs
(sinks)
• Temperature and heat energy are also defined by the
heat flow in and out of the spacecraft material(s), and
by the thermal heat capacity of the material(s)
– Heat capacity is the heat energy required to change the
temperature of a material by one degree
• Temperatures of specific areas of a spacecraft are
also affected by the conductivity of other components
in contact, and by the other heat transfer
mechanisms, not just surface radiation absorption
and emission
Thermal Control Systems
Thermal Control Systems
Heat sources (inputs)
1. Solar radiation (insolation) at the Earth is 1340 W/m2
(seasonal average)
• Actual value ranges from approximately 1310 to 1420 W/m2
• The equivalent (peak blackbody) temperature of the Sun is
approximately 5,800K
2. Earth's infrared radiation is 240 W/m2 in LEO from the
atmosphere and surface heat (varies with orbit altitude)
• The Earth's equivalent blackbody temperature is
approximately 290K
3. Earth's reflected solar energy is 400 W/m2 (30% of direct
solar energy reflected in the upper atmosphere) (varies with
orbit altitude)
4. Spacecraft internal heat is also a heat source that comes
from electrical power used in nearly all onboard systems
Thermal Control Systems
Heat sinks (outputs)
• Heat removal from spacecraft is almost always done
by radiated energy (thermal radiators)
– Several early manned missions used evaporative
cooling, but using liquids for cooling is inefficient
• Heat emission to space is based on the infrared
blackbody temperature of the spacecraft surface(s)
(emission component) and the temperature of the
background (absorption component)
Thermal Control Systems
A quick look at the challenges of a spacecraft thermal system
Spacecraft temperatures vary widely since:
• Spacecraft encounter temperature extremes that range
from the orbit of Mercury to the orbit of Pluto and beyond
• Each spacecraft has a variety of heat inputs and heat
outputs
• Each spacecraft has complex surface geometries and
materials which affect emission and absorption
• There are various levels of heat transfer inside and at the
surface of the spacecraft that change almost continually
– Transfer includes complicated paths of conduction and
radiation
• There are continual changes in heat inputs and heat
outputs because of spacecraft movement, orbital motion,
and/or flight trajectories
Thermal Control Systems
Simplified spacecraft thermal system design
Spacecraft thermal control system design is a difficult processes because
of the interplay of heat transfer among the components, and because of
the extreme temperatures encountered in space, and because of the everchanging heat inputs and outputs during the lifetime of a spacecraft. A
simplified sequence in a spacecraft thermal control system design might
look like the following (Griffin & French):
1.
2.
Spacecraft mission specified
Space environment (heat characterization) specified by mission
trajectory
3. Spacecraft component selection made for mission objectives
4. Temperature requirements established for components
5. Structural assemblies established
6. Surface temperatures approximated
7. Heating or cooling needs approximated for all onboard systems
8. Passive thermal elements selected
9. Thermal range reexamined for critical components
10. Active systems adopted to satisfy temperature range requirements
11. Models used to verify overall thermal design
12. Tests on thermal models made in vacuum chamber with controlled
thermal environment
Thermal Control Systems
Basic thermal system considerations
• A spacecraft thermal system design is far beyond the
scope of these lectures, but temperature control
considerations for extreme conditions in space can
be outlined with a simplified approach
Thermal Control Systems
• Extreme temperatures in space demand thermal
regulation to moderate the temperature range
experienced by the outer spacecraft structure, with an
even narrower temperature range within the sensitive
systems and subsystem components
• Critical electronics and sensors may require an even
smaller temperature range for reliable operation
• Cabin environments for crews are even more
restrictive, with requirements typically 68-80oF
Thermal Control Systems
Earth-orbit temperature range approximately -250oF to
+300oF depending on surface absorptivity and
emissivity (and thermal conductivity, specific heat, and
other factors)
Temperature range limit s (for this example)
• For a spacecraft interior use -120oF to +200oF
• For interior modules use -60oF to +100oF
• For critical components use +65oF to +85oF
Thermal Control Systems
This is a graphic representation of the extreme temperatures encountered in the nearEarth space environment, and the temperature range restrictions on a hypothetical
spacecraft
Passive thermal components are capable of reducing the temperature extremes, but
only to a limited range
Beyond this, active systems would be required for a narrower temperature range
Thermal Control Systems
Passive thermal components
1. Radiator panels
2. Coatings
3. Heat pipes
4. Insulation
5. Conductive structures and components
6. Louvers
7. Sun shields
8. Radioisotope heater units
Thermal Control Systems
1. Radiator panels
• Radiator panels are used as passive heat radiators for
moderate heat sources on spacecraft
• Higher heat loads can also be passed to space
through radiator panels, but the heat is transferred
with liquid cooling loops, and pumped through the
panels, with larger panes for higher heat removal
requirements
Thermal Control Systems
2. Coatings
• The use of a surface coating is based on the surface
material's emissivity and absorptivity
• Although Kirchkoff's law states that good emitter is
also a good absorber, few materials follow this law
• The following table shows that only black paint and
polished metal follow the equivalency of absorption
and emission
• Coating are one of the most common passive thermal
control element used on spacecraft.
Surface Finish
Absorptivity (Beginning-of-Life)
Emissivity
Thermal Control Systems
Optical Solar Reflectors
8 mil Quartz
Mirrors
.05 to .08
.80
2 mil silvered
Teflon
5 mil Silvered
Teflon
2 mil Aluminized Teflon
.05 to .09
.05 to .09
.10 to .16
.66
.78
.66
5 mil Aluminized
.10 to .16
.78
.20 to .25
.17 to .20
.18 to .20
.22 to .28
.85
.92
.91
.88
.92 to .98
~.97
.89
.84
.34
.38
.41
.46
.55
.67
.75
.86
.08 to .17
.09 to .17
.19 to .30
.25 to .86
.04
.03 to .10
.03
.04 to .88
(Material degrades in sunlight)
.32
~.22
.9
.34
.86
.80
.1
Teflon
White paints
S13G-LO
Z93
ZOT
Chemglaze A276
Black Paints
Chemglaze Z306
3M Black Velvet
Aluminized Kapton
1/2 mil
1 mil
2mil
5 mil
Metallic
Vapor Deposit Aluminum (VDA)
Bare Aluminum
Vapor Deposit Gold
Anodized Aluminum
Mylar
1/4 mil Aluminized Mylar, Mylar side
Beta Cloth
Astro Quartz
MAXORB
Thermal Control Systems
3.Heat pipes
• The heat pipe is a combined heat conduction and
phase change mechanism which has very efficient
heat conductivity properties
• No energy is required for operation, although
mechanical systems can be used to augment the heat
transfer capability of the heat pipe
Thermal Control Systems
4. Insulation
• Insulation is effective for both heat rejection and heat
retention
• The typical multilayer insulation blanket reduces the
heat flow in or out by restricting the heat flow
• The blanket layers are designed specifically for the
thermal conditions associated with the spacecraft
(internal heat) and the space environment (external
heat)
Thermal Control Systems
5. Conductive structures and components
• The heat conductivity of the spacecraft structures
and assemblies plays an important role in the thermal
design of the spacecraft
• Temperature critical components are placed on the
structural bus in a location where the heat
input/output can be controlled within a specified
range, either passively (preferably - less energy
needed) or actively
• A node analysis is made of the structural assembly
with the heat sources and sinks estimated to verify
the temperature range expected for the components
Thermal Control Systems
6. Louvers
• These passive elements can provide both heating and
cooling capability by adjusting the opening angle for
increased thermal shading (cooling), or increased
heat exposure (heating)
Thermal Control Systems
7. Sun shields
• The sun shield restricts the heat input of a spacecraft
with either a fixed or movable shield (louver, for
example)
• Also used also to reduce the light exposure from the
Sun for sensitive instruments (Hubble Space
Telescope, Infrared Astronomical Satellite, Galileo for
VEEGA)
Thermal Control Systems
8. Radioisotope heater units
• Radioisotope heater units (RHUs) are small devices
that produce heat by the radioactive decay of Pu-238
• The small units generate about one watt of heat from
the decay of a few grams of Pu-238 for several
decades (Pu-238 has a 88 year half-life)
• Spacecraft RHUs are used to heat critical
components and subsystems, and reduce spacecraft
complexity for the thermal subsystems, and help
reduce the overall complexity of a spacecraft
• The Cassini spacecraft at Saturn employs 82 of the
RHUs
Thermal Control Systems
Thermal Control Systems
Active thermal components
1.
2.
3.
4.
5.
Electric cooling devices
Stored cryogenics
Liquid heating/cooling loops
Electric heaters
Shutters
Thermal Control Systems
1. Electric cooling devices
1. Stirling cycle coolers
Active coolers based on causing a working gas to undergo a Stirling cycle
consisting of two constant volume processes and two isothermal
processes. Devices consist of a compressor pump and a displacer unit
with a regenerative heat exchanger. Stirling cycle coolers were the first
active cooler to be used successfully in space and have proved to be
reliable and efficient.
2. Pulse tube coolers
Pulse tube coolers are similar to the Stirling cycle coolers with a different
thermodynamic processes, and consist of a compressor and a fixed
regenerator. There are no moving parts and reliability is theoretically higher
than Stirling cycle machines.
3. Joule-Thompson (J-T) coolers
These coolers work using the Joule-Thomson (Joule-Kelvin), effect which
occurs as a gas is forced through a thermally isolated porous plug or
throttle valve by a mechanical compressor unit leading to isenthalpic
cooling. Although this is an irreversible process, with correspondingly low
efficiency, J-T coolers are simple, reliable, and have low electrical and
mechanical noise levels. A J-T stage driven by a valved linear compressor
is on the planned Planck CMBR telescope
Thermal Control Systems
1. Electric cooling devices
4. Sorption
Sorption coolers are essentially J-T coolers which use a thermochemical process to provide gas compression with no moving
parts. Powdered sorbent materials (e.g. metal hydrides), are
electrically heated and cooled to pressurize, circulate, and adsorb a
working fluid such as hydrogen.
5. Reverse Brayton
Reverse/Turbo Brayton coolers have high efficiencies and are
practically vibration free. Coolers consist of a rotary compressor, a
rotary turbo-alternator (expander), and a counterflow heat
exchanger. The compressor and expander use high-speed
miniature turbines on gas bearings and small machines are thus
very difficult to build. They are primarily useful for low temperature
experiments.
6. Adiabatic demagnetization refrigeration (ADR)
Adiabatic demagnetization refrigeration has been used on the
ground for many years to achieve milli-Kelvin temperatures after a
first stage cooling process. The process utilizes the magnetocaloric effect with a paramagnetic salt.
Thermal Control Systems
1. Electric cooling devices
7. 3He coolers
In addition to its use as a stored cryogen the properties of can
be used to achieve temperatures below 1K with closed cycle
sorption coolers and dilution refrigerators.
8. Optical cooling
In recent years the principle of optical cooling has been
developed and demonstrated. The principle of anti-Stokes
fluorescence in ytterbium doped zirconium fluoride is used to
provide vibration-free solid-state cooling.
9. Peltier effect coolers
Solid-state Peltier coolers, or thermo-electric converters, are
routinely used in space to achieve temperatures above 170K
(e.g. the freezers aboard the International Space Station). These
devices work on the same principle as the Seebeck effect, but in
reverse.
Thermal Control Systems
Electric cooling devices – devices and approximate temp range
Thermal Control Systems
2. Stored cryogenics
• Dewars containing a cryogenic liquid such as liquid
helium or solid neon may be used to achieve
temperatures below those offered by radiators
• These systems provide excellent temperature stability
with no exported vibrations but substantially increase
the launch mass of the vehicle and limit the lifetime of
the mission to the amount of cryogen stored
• They have also proved to be of limited reliability
Thermal Control Systems
3. Liquid heating/cooling loops
• Refrigeration loops are used primarily for cooling but
can also be used for heating components with the
resulting waste heat within the system
• These cooling loop is generally a liquid/gas phase
process because of the large heat content in the
phase change for many materials
• The process transfers heat through a heat exchange
system and requires a radiator, compressor and
liquid pump(s) for operation
Thermal Control Systems
4. Electric heaters
• Resistive heating is relatively simple and efficient for
most spacecraft thermal control applications
• Temperature range can be maintained with a
thermocouple/thermostat, although ranges less than
approximately ±2oC can create control problems
• Components for the resistive heater are simple: a
resistor, electric current, and a temperature control
limiter (thermostat)
Thermal Control Systems
5. Shutters
• Active (powered) louvers are shutters and serve the
same purpose as the louver - heat shading or
exposure
• Shutters are generally for larger heat loads and/or
more responsive thermal control
Thermal Control Systems
The Mercury capsule is used as a spacecraft thermal
system example because of its simplicity relative to the
crew reentry vehicles developed later, and because of
the temperature extremes of this and any other reentry
capsule
• A narrow range of temperatures within the cabin and
even narrower in the crew member's flight suit further
restricted the internal heat loading range
• Durable, high temperature, passive elements were
required to reduce the nearly 3,000oF reentry shield
temperature to several hundred degrees at the
capsule's internal structure
Thermal Control Systems
Active thermal system components allowed internal and external
heat loads to be redistributed, providing a modest thermal
environment for the astronaut during all flight phases
Cabin temperatures for the MA-6 flight shown below (NASA)
Thermal Control Systems
• Mercury capsule structure
was primarily titanium
sheets (0.01"), seam
welded, and stiffened with
titanium longitudinal
stringers (0.25") at 15o
positions around the
conical capsule
• Bulkheads for the internal
pressure vessel were
titanium sheet with added
hoops and rings for
increased strength
Thermal Control Systems
• The capsule reentry heat
shield was constructed
of an ablation shield,
nickel and beryllium
shingles, durable
insulation, and a
titanium shell structure
on the inner and outer
walls
• Mercury’s ablation shield
was composed of glass
fiber and resin which
vaporized during
maximum heating at
reentry
Thermal Control Systems
• Effective cooling was produced with only 2-4 lb of
material vaporized during reentry
Temperatures encountered in and around the MA-6
capsule shown below (exit = reentry phase exit) (temps
in oF)
Thermal Control Systems
•
Heating model for Mercury cabin temperatures. Plot shows the variable
heating on orbit due to exposure to the Sun and internal heat load during
orbit and reentry. Beta is the orbital solar reference angle which is
approximated by the orbital inclination (McDonnell).
Thermal Control Systems
Beta angle is the angle between the Sun vector and the orbit plane of
an Earth-orbiting object
• High beta = approximated by high inclination orbit with greater
time in sunlight and less time in eclipse (greater solar heating)
• Low beta = approximated by low inclination orbit with less time in
sunlight and more time in eclipse (less solar heating)
• Precession of orbit plane introduces a progressive change in beta
angle (applies to almost all orbits)
• Solar beta angle cutout – launch restriction for the Space Shuttle
orbiters docked to the ISS because of a high beta angle and
possible heat overload during the mission
Thermal Control Systems
Reentry angle, reentry velocity and heat loading shown for a capsule model.
QAVE is heat load average, and QMAX is the maximum heating for entry
parameters (reentry angle γ and entry velocity for suborbital flight)
(McDonnell).
Thermal Control Systems
Heat shielding
• Heat shields are solid surfaces that protect reentry
capsules from the extreme heat of atmospheric friction
at hypersonic, suborbital, orbital speeds, and lunar and
Mars return speeds
• Heat shields are generally a blunt curved plate
(Mercury, Gemini, Apollo), or a blunted cone (missile
warheads, Galileo probe)
• Highest temperatures found in the friction layers and
reentry surfaces are ahead of the shield’s leading
surface at the shock front
Thermal Control Systems
Heat shielding
• Energy dissipated is same as kinetic energy of capsule
= ½ mV2
• As velocity increases, temperature should increase as
kinetic energy to an exponential power (most of the
kinetic energy is dissipated by shock wave)
• Actual temperature increase using a coincidental ruleof-thumb is Tmax (Kelvin) = entry velocity in m/s
– 7,800 m/s entry ~ 7,800 K (shock layer maximum
temp)
– Smaller than expected max shock temp is due to
heated gas specific energy increase with increasing
temp
Thermal Control Systems
Heat shielding
• Most common type of heat shielding for crew capsules
is an ablative tile cover that also provides heat
insulation for the capsule interior
• Ablation, or burn-off, occurs in the outer layer as the
material melts, chars, and sublimates
• Inner layer pyrolizes (heated chemical change without
oxygen) and generates gas that pushes outward,
providing a cooler boundary layer on the shield
• Carbon expelled into the shock layer also reduced the
temp at heat shield because of optical and IR opacity
(opaque layer blocks radiation heating of the shield
from the high-temp shock layer)
Thermal Control Systems
Heat shielding
• Mercury capsule heat shield was
a fiberglass and phenolic resin
within aluminum honeycomb
structure
• After reentry heating dropped
below ablation temps, the heat
shield was jettisoned to prevent
shield heat from being
transferred to the capsule
• Mercury capsule landing airbag
was then deployed to soften the
landing
Thermal Control Systems
Heat shielding
• Other ablative heat shields include the phenolic epoxy
resin and silica fiber used on Apollo capsules called
Avcoat that is being used on NASA’s new Orion Crew
Exploration Vehicle (CEV)
• Phenolic Impregnated Carbon Ablator (PICA) was
developed and used on robotic spacecraft including
NASA’s Stardust sample return mission from comet
Wild 2
• SpaceX’s Dragon capsule uses a proprietary variation
of the PICA material named PCIA-X that is capable of
protecting the capsule returning from the Moon or Mars
Thermal Control Systems
Heat insulation
• Thermal protection for reentry vehicles also includes
durable insulation tiles that do not burn off (ablate)
• Metallic heat shielding on the Mercury and Gemini
capsule’s aft and mid sections was replaced with
lighter-weight, high-temperature insulation layers on
the Apollo capsule
• More dramatic changes in thermal protection for
reentry vehicles was made in the design of the Space
Shuttle Orbiter’s passive Thermal Control System (TPS)
tiles
Orbiter Passive Thermal Protection System
Orbiter TPS original design requirements
• Limit aluminum structure temperature to 350 °F
• 100 mission capability with cost-effective unscheduled
maintenance/replacement
• Withstand surface temperatures from -250° to 2,800 °F
• Maintain the moldlines for aero and aero-thermo
requirements
• Attach to aluminum structure
• Economical weight and cost
Thermal Control Systems
Orbiter passive thermal tile types
•
Reinforced Carbon-Carbon (RCC) - Used on the nose cap and wing
leading edges where reentry temperatures exceed 1,260° C (2,300° F)
•
High-temperature Reusable Surface Insulation (HRSI) - Used primarily
on the Orbiter belly where reentry temperatures are below 1,260° C
•
Toughened Unipiece Fibrous Insulation (TUFI) - A stronger, more
durable tile that is replacing high and low temperature tiles in highabrasion areas
•
Low-temperature Reusable Surface Insulation (LRSI) - Originally used
on the upper fuselage, but now mostly replaced by AFRSI
•
Advanced Flexible Reusable Surface Insulation (AFRSI) - Quilted,
flexible surface insulation blankets used where reentry temperatures
are below 649° C (1,200° F)
•
Fibrous Refractory Composite Insulation (FRCI) - FRCI tiles that have
replaced some of the HRSI 22 lb tiles provide improved strength,
durability, resistance to coating cracking
•
Felt reusable surface insulation (FRSI) - Nomex felt blankets that are
used on the upper regions of the Orbiter where temperatures are
below 371° C (700° F)
TPS Surfaces
Lower Surface
Upper Surface
TPS Legend
HRSI (Black) Tiles
LRSI (White) Tiles
AFRSI Blankets
FRSI
RCC
Glass
Exposed Metallic Surfaces
Side View
Leading Edge RCC
Fixed Upstream Gap Between
Panel and Tee Seal
E
A
A
E
HRSI Tiles
Variable Downstream Gap
Between Panel and Tee
Seal For Thermal Expansion
Allowance
Section A-A
B
B
Detail D
Interface Gap
Between
RCC and
HRSI Tiles
D
Section C-C
Section B-B
Upper LESS
Access Panel
HRSI Tiles
RCC
Tee Seal Web
(In Background)
Thermal Barrier
I nconel
Attachments
C
Upper
Left
Wing
Frank Jones NASA, KSC
C
RCC Panel
(In Foreground)
Upper Wing AFRSI
ACSS Hardware
I nconel
Insulators
Wing Spar
Lower Wing HRSI Tile
Horsecollar
Peripheral Gap Filler
Section E-E
Lower LESS Access
Panel HRSI Tile
TPS - HRSI
High-temperature Reusable Surface Insulation tiles are
used to insulate the Orbiter's underside aluminum
alloy structure from the reentry heat that ranges
from -157oC to 1,260° C (-250oF to 2,300° F)
Like the other rigid Orbiter insulation tiles, the HRSI
tiles were designed withstand:
– On-orbit cold soaking
– Repeated thermal shock from heating and cooling
– Extreme acoustic environment during launch and
reentry reached 165 decibels
HRSI Tile Configuration
HRSI Tiles - Black RCG Coating
LRSI Tile White
Glass Coating
Gap
Step
Densified IML Surface
Koropon-Primed
Structure
Silicone RTV Adhesive
SIP
Uncoated Tile
Filler Bar
Coating Terminator
Frank Jones NASA, KSC
Wing Tiles – Discovery (STS-114)
HRSI Tiles - Black RCG Coating
LRSI Tile White
Glass Coating
Gap
Step
Densified IML Surface
Koropon-Primed
Structure
Silicone RTV Adhesive
SIP
Uncoated Tile
Filler Bar
Coating Terminator
Thermal Control Systems
Reentry heat shielding – other technology
• Other than ablative shielding on reentry capsules and
the passive-radiative insulation used on the Space
Shuttle Orbiters, other reentry heat protection
technologies include radiatively cooled ceramic
materials and exotic metal alloys
• Protective heat shielding technology for reentry and
hypersonic vehicles is also changing because of the
improvement in the heat range of advanced structural
materials
Thermal Control Systems
Traditional and newer structural materials
Thermal Control Systems
Thermal insolation - Aerogel
• Aerogel is a synthetic, porous, ultralight material
derived from a gel in which the liquid component of the
gel has been replaced with a gas by supercritical drying
• The result is a solid with extremely low density and
thermal conductivity used in a variety of applications
including several exploration spacecraft
• Nicknames include "frozen smoke", "solid smoke",
"solid air" or "blue smoke" owing to its translucent
nature and the way light scatters in the material
(Rayleigh scattering from fine structure)
Thermal Control Systems
Solid Aerogel block
Thermal Control Systems
Aerogel insulation
Thermal Control Systems
Aerogel strength
Thermal Control Systems
Thermal insolation - Aerogel
• Aerogel materials include
silica (silica dioxide), carbon,
alumina (aluminum oxide),
and other metals
• Aerogel was used for the
particle collector “catchers
mitt” on NASA’s Stardust
sample return spacecraft
Thermal Control Systems
References
• Aherns, D. C., Meteorology Today, 1991, West Publishing, NY
• Griffin & French, J., R., Space Vehicle Design, 1991, AIAA,
Washington, D.C.
• Gilmore, David, Satellite Thermal Control Handbook, 1994,
The Aerospace Corporation Press, CA
• Larson & Wertz (ed.), Space Mission Analysis and Design,
1992, 2nd edition, Kluwer Academic Publishers, Boston
• NASA, First United States Manned Orbital Space Flight,
February 1962
• McDonnell Aircraft, Manned Satellite Capsule: Vol 2 Technical Proposal, 1958