Thermal Control Systems

Thermal Control Systems
SP 300

Definitions Thermal Space Environment Passive and Active Subsystems Spacecraft Examples

Definitions

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 = K (boiling water K) Celsius – 0o Celsius is freezing point of water, 100oC is boiling point of water, oC is absolute zero Fahrenheit – +32oF is freezing point of water, 212oF is boiling point of water, °F is absolute zero

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

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

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

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)

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

Absorptivity α Emissivity ε

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

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

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

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 Space Environment

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

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) Near Earth (1.0 AU) 400 K (260oF) Near Mars (1.5 AU) 300 K (80oF) 20 K (-280oF) Near Jupiter (5.3 AU) 100 K (-280oF) 20 K (-420oF) Near Pluto (40 AU)

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

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

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)

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

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 ever-changing 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): Spacecraft mission specified Space environment (heat characterization) specified by mission trajectory Spacecraft component selection made for mission objectives Temperature requirements established for components Structural assemblies established Surface temperatures approximated Heating or cooling needs approximated for all onboard systems Passive thermal elements selected Thermal range reexamined for critical components Active systems adopted to satisfy temperature range requirements Models used to verify overall thermal design Tests on thermal models made in vacuum chamber with controlled thermal environment

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

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

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

This is a graphic representation of the extreme temperatures encountered in the near-Earth 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

Passive and Active Thermal Systems

Passive thermal components Radiator panels Coatings Heat pipes Insulation Conductive structures and components Louvers Sun shields Radioisotope heater units

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

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 Optical Solar Reflectors 8 mil Quartz Mirrors .05 to .08 .80 2 mil silvered Teflon .05 to .09 .66 5 mil Silvered Teflon .78 2 mil Aluminized Teflon .10 to .16 5 mil Aluminized Teflon White paints S13G-LO .20 to .25 .85 Z93 .17 to .20 .92 ZOT .18 to .20 .91 Chemglaze A276 .22 to .28 .88 Black Paints Chemglaze Z306 .92 to .98 .89 3M Black Velvet ~.97 .84 Aluminized Kapton 1/2 mil .34 .55 1 mil .38 .67 2mil .41 .75 5 mil .46 .86 Metallic Vapor Deposit Aluminum (VDA) .08 to .17 .04 Bare Aluminum .09 to .17 .03 to .10 Vapor Deposit Gold .19 to .30 .03 Anodized Aluminum .25 to .86 .04 to .88 Mylar 1/4 mil Aluminized Mylar, Mylar side (Material degrades in sunlight) Beta Cloth .32 Astro Quartz ~.22 MAXORB .9 .1 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

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)

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

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)

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)

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

Active thermal components Electric cooling devices Stored cryogenics Liquid heating/cooling loops Electric heaters Shutters

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 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

1. Electric cooling devices 4. Sorption Sorption coolers are essentially J-T coolers which use a thermo-chemical 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 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 magneto-caloric effect with a paramagnetic salt.

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.

Electric cooling devices – devices and approximate temp range

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

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

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)

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

Spacecraft Examples

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

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)

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

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

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)

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).

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

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).

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

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 rule-of-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

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)

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

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

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

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 high-abrasion 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 Side View
HRSI (Black) Tiles LRSI (White) Tiles AFRSI Blankets Side View FRSI RCC Glass Exposed Metallic Surfaces

Leading Edge RCC Frank Jones NASA, KSC A Fixed Upstream Gap Between
Panel and Tee Seal A A E E HRSI Tiles Section A-A Variable Downstream Gap Between Panel and Tee Seal For Thermal Expansion Allowance Upper Wing B B Detail D Interface Gap Between RCC and HRSI Tiles Section B-B Upper LESS Access Panel HRSI Tiles Upper Wing AFRSI D RCC Tee Seal Web (In Background) Thermal Barrier Section C-C I nconel Attachments C ACSS Hardware C I nconel Insulators Upper Left Wing Wing Spar RCC Panel (In Foreground) Lower Wing HRSI Tile Horsecollar Peripheral Gap Filler Frank Jones NASA, KSC Lower LESS Access Panel HRSI Tile Section E-E

Repeated thermal shock from heating and cooling
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 Gap LRSI Tile White Glass Coating 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 Gap LRSI Tile White Glass Coating Step Densified IML Surface Koropon-Primed Structure Silicone RTV Adhesive SIP Uncoated Tile Filler Bar Coating Terminator

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

Traditional and newer structural materials

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)

Solid Aerogel block

Aerogel insulation

Aerogel strength

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

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

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