Structures / Thermal Group

Structures/Environmental 

Protection Team 

Keiichi Narumi 

Ben Hochman 

Christian Hirschen 

Raymond Chow 

Jennifer Stothers 

Vivek Chetty 

Scott Wong

Overview of Our Task 

Structural Expectation 

1. Space 

2. Re-entry Re entry 

3. Landing on Mars 

Past Structural Failure 

Thermal/Radiation Control on Board 

Thermal/Structural Protection Against Re- Re 

Entry/Environment 

Comparison of Each Option of Structural 

and Landing Materials

Environment in Space and on Mars 

Why do we have to take the environmental conditions into account? 

Constraints for thermal and radiation shielding and the requirements 

for the materials which have to be used (e.g. entry in atmosphere 

causes high temperatures) 

Determine the life cycle of all the equipment (e.g. radiation causes 

malfunction) 

Temperature limits for spacecraft electronics (e.g. batteries 

from 5 to 20 deg C) 

Structures and environmental Protection against Thermal/Radiation in Transit and on Ground

1. Thermal Conditions 

•in Space 


– Space is more or less a vacuum except the constituents of the 

exosphere hydrogen and helium in very low densities 

– Temperature depends mainly on whether you are exposed or not to 

sunlight and how far you are away from the sun 

– The coldest temperature is 3K, this is the temperature of the cosmic 

microwave background radiation (CMB), so without any other heat 

sources, gas will come in equilibrium with the CMB 

– There is not one temperature for the space, but it is very 

variable 



Earth Atmosphere 

http://liftoff.msfc.nasa.gov/academy/space/atmosphere.html 



• on Mars 


– The six most common components of the atmosphere are : 

•Carbon Dioxide (CO2): 95.32% 

•Nitrogen (N2): 2.7% 

•Argon (Ar): 1.6% 

•Oxygen (O2): 0.13% 

•Water (H2O): 0.03% 

•Neon (Ne): 0.00025 % 

– Martian air contains only about 1/1,000 as much water as our air 

Enough to condense out and form clouds and local patches of 

morning fog 



• on Mars 


– The average recorded temperature on Mars is -63° C (-81° F) with a 

maximum temperature of 20° C (68° F) and a minimum of -140° C (-220° F) 

– During entry into the Martian atmosphere the temperature can reach values 

up to 2000 deg C, even though the density is much smaller than on earth 

– For the different mars missions the pressure varied between 6 and 11 

millibars, which is around 1/1000 of the earth pressure 



http://www.dfg.de/raumtransportsysteme/rueckkehr_2.html 



Mars Statistics 

Mass (kg) 6.42E+23 

Mass (Earth = 1) 1.07E-01 

Equatorial radius (km) 3,397.20 

Equatorial radius (Earth = 1) 5.33E-01 

Mean density (gm/cm^3) 3.94 

Mean distance from the Sun (km) 227,940,000 

Mean distance from the Sun (Earth = 1) 1.5237 

Rotational period (hours) 24.6229 

Rotational period (days) 1.025957 

Orbital period (days) 686.98 

Mean orbital velocity (km/sec) 24.13 

Orbital eccentricity 0.0934 

Tilt of axis (degrees) 25.19 

Orbital inclination (degrees) 1.85 

Equatorial escape velocity (km/sec) 5.02 

V isual geometric albedo 0.15 

Magnitude (Vo) -2.01 

Minimum surface temperature -140°C 

Mean surface temperature -63°C 

Maximum surface temperature 20°C 

Atmospheric pressure (bars) 0.007 

www.solarviews.com/eng/mars.htm 



Martian Atmosphere 

http://abenteuer-universum.vol4u.de/mars.html 



Temperature Limits 

• Spacecraft electronics from around 0 to 40 deg C 

• Batteries from 0 to 15 deg C 

• Silicon Solar Cells from -100 to 100 deg C 

• Structure from -45 to 65 deg C 

• Structures supporting equipments with extremely accurate appointing 

requirements are limited to +/- 0.5 deg C variation 


2. Radiation 


• Three different types of radiation 

– SPE: Solar Particle Elements 

– GCR: Galactic Cosmic Ray 

– Trapped Radiation 



– SPE: Solar Particle Elements 

• Produced by the acceleration of solar plasma by strong electromotive 

forces and produced by acceleration across the transition shock 

boundary of propagating coronal mass ejecta 

• Occur in association with solar flare 

• Are rapid increases in the flux of energetic particles (~1 MeV to above 

1GeV), lasting from several hours to several days 

• Degrade solar panels, increase background noise and cause illnesses 



– GCR: Galactic Cosmic Ray 

• GCR: galactic origin 

• A single particle can cause a malfunction in common electronic 

components → single event phenomena (SEP) 

• Can also generate background noise 

• Create spurious which masquerade as real signals 

• No full understanding of the physical interaction of GCR with shielding 

and body tissues available yet 

→ large impact on costs 



– Trapped Radiation (Van Allen radiation belts) 

• In this case particles are trapped within the confines of the geomagnetic 

field 

• The trapped radiations consist mainly of protons and electrons reaching 

maximum intensity at an altitude 3600km followed by a minimum at 

7000 and a second very broad maximum at 10000km 

• Having energies greater than 30keV 

• The accumulation of these radiation events can cause permanent damage to 

individual detectors, electronic devices, solar arrays and sensors 


Thermal Environment 

Direct Sunlight 

Radiation 

Sunlight Reflected 

by Planets (albedo ( albedo) 

Infrared (IR) 

Radiation Emitted 

from a Planet’s Planet s 

Atmosphere or 

Surface 

Surface http://www-istp.gsfc.nasa.gov/istp/outreach/images/Solar/Events/skylab.jpg

Thermal Control 

Heat Emitted as IR Radiation from the 

Spacecraft must be Balanced with the 

Heat Dissipated by Internal Components 

and Heat Absorbed from the Environment

Custom Tailored 

A generic thermal control system would be 

heavy and expensive 

It is most practical to design a thermal 

control system for the specific needs of 

the mission while taking into account the 

worst case (hot & cold) of the 

environment as well as the allowable 

temperatures of the components and 

cargo.

Direct Solar Radiation 

Between planets, 

external heat 

comes from direct 

sunlight which falls 

off as the square of 

the distance to the 

sun. 

Solar constant: 

-Venus Venus 2700W/m 2 

-Saturn Saturn 15W/m 2

Heat Transfer Issues 

As the density of the spacecraft increases 

not all components can be mounted 

directly to a thermal control system and 

must be stored internally 

Heat transfer systems such as heat pipes 

must be used for internally mounted cargo

The Hot & Cold of Space 

Components such as batteries require 

small variations in temperatures (0 to 

15°C) 15 

Solar Panels can withstand much larger 

variations in temperature (-150 ( 150 to 110°C) 110 

Fuel tanks and lines must remain above 

freezing (15 to 40°C) 

40

Cryogenic Cooling 

IR detectors often require continuous 

extreme cold (125 K and below) 

Expendable fluids or gasses can be 

used as stored cryogens to absorb 

heat but are difficult to store due to 

volume constraints of the spacecraft 

Refrigeration systems require large 

amount of electrical power

Heat Balance 

Careful heat balance estimations must be 

made and although a wealth of 

information is available for earth orbiting 

spacecraft, much less is known for 

interplanetary missions 

For much longer missions, nuclear power 

supplies could be used and the waste heat 

distributed around the spacecraft

Structural Failure 

Example: Columbia-February Columbia February 1, 2003 

-One One large piece and at least two smaller pieces of 

insulating foam separated from the External Tank left 

bipod ramp area. 

-Later Later analysis showed that the larger piece struck 

Columbia on the underside of the wing, around 

Reinforced Carbon-Carbon Carbon Carbon panels 5 through 9. 

-The The foam piece was approximately 21 to 27 inches 

long and 12 to 18 inches wide, tumbling at a 

minimum of 18 times per second and 416 to 473 

mph at time of impact. 

Columbia Accident Investigation Board

Structural Design Philosophy and 

Criteria 

Launch loads are affected by 

acoustics 

acoustics 

engine vibrations 

air turbulence 

gusts

-Use Use a design-allowable design allowable strength for the selected 

material that we expect 99% of all specimens will 

equal or exceed. 

-From From available environmental data, derive a design 

limit load equal to the mean value plus three 

standard deviations. 

-Multiply Multiply the design limit load by a factor of safety, 

and then show that the stress level at this load does 

not exceed the corresponding allowable strength. 

-Test Test the structure to verify design integrity and/or 

workmanship, to correlate analytical models, and to 

protect against human errors.

Terms 

Load Factor 

Limit Load 

(or design limit load) 

Allowable Load or 

Stress 

Factor of Safety, FS 

Design Load 

Design Stress 

Margin of Safety, 

MS 

A multiple of weight on Earth, representing the 

force of inertia that resists acceleration. 

The maximum load expected during the mission 

or for a given event, at a specified or selected 

statistical probability (typically 99% for 

expendable launch vehicles and 99.87% for 

launches with humans aboard). 

The highest load or stress a structure or material 

can withstand without failure, based on statistical 

probability (usually 99%; i.e., only 1% chance the 

actual strength is less than the allowable). 

A factor applied to the limit load to obtain the 

design load for the purpose of decreasing the 

chance of failure. 

Limit load multiplied by the yield or ultimate 

factor of safety; this value must be no greater than 

the corresponding allowable load. 

Predicted stress caused by the design load; this 

value must not exceed the corresponding 

allowable stress. 

A measure of reserve strength: 

design criteria 

Definition 

Allowableload 

(or stress) 

MS = 

−1 

≥ 0 

Design load (or stress) 

to satisfy 

Table 11-53. Terms and Criteria Used in Strength Analysis

4. No structural test 

Options 

1. Ultimate test of dedicated qualification article (1.25 x limit) 

2. Proof test of all flight structure (1.1 x limit) 

3. Proof test of one flight unit of a fleet (1.25 x limit) 

Design Factors of 

Safety 

Yield 

1.0 

1.1 

1.25 

1.6 

Ultimate 

1.25 

1.25 

1.4 

2.0 

Table 11-54. Typical Test Options and Factors of Safety for Missions without Humans Aboard.

Preliminary Sizing of Structural Members 

Stiffness 

Finite element model 

Stiffness 

Strength 

Statically determinate structure - Free-body Free body diagrams 

Statically indeterminate structure – Finite element 

analysis 

Strength 

Weight 

Compare a component’s component s weight with its allocation 

Best design ≠ lightest design 

Weight

History of Failed Mars Missions 

Marsnik 1 & 2 – 10/10/60 and 10/14/60 

Its launch vehicle did not produce enough thrust to put the probe probe 

into the 

proper trajectory for Mars, and it fell back to Earth. Marsnik 2 was 

identical to Marsnik 1. The third stage of its rocket also failed failed 

during 

launch four days later. 

Sputnik 29 – 10/24/62 

Sputnik 29 failed to reach Earth orbit and broke apart when the Molnya 

rocket's fourth-stage fourth stage engine failed. 

Mars 1 – 11/1/62 

Mars 1’s 1 s signal was lost about half way through the journey on March 21, 21, 

1963. 

Sputnik 31 – 11/4/62 

Sputnik 31 reached Earth orbit but failed while firing the rocket rocket 

motor that 

would have put it on a trajectory for Mars. 

Mariner 8 – 5/8/71 

The main Centaur engine ignited 265 seconds after launch but the upper 

stage began to oscillate in pitch and tumble out of control.

Mars 2 – 5 /19/71 

The landers were released from the orbiter four and a half hours 

before the spacecraft entered Martian orbit on 11/27/1971. 

Lander was destroyed on impact. 

Mars 4-7 4 7 – 7/21/73 to 8/9/73 

A decision had been made two years prior to launch to replace the the 

gold components in the transistors with aluminum in order to save save 

Soviet gold resources. Mars 4 flew past Mars. Mars 5 lasted a 

few days. Mars 6 crashed and burned due to fast impact. Mars 7 

missed the planet. 

Phobos 1-2 1 2 – 7/7/88, 7/12/88 

Phobos 1: Battery died because it lost the ability to point its solar 

arrays toward the sun. 

Phobos 2: Turn antenna away from Earth to point the camera 

toward Phobos. Never turn back around. 

Mars Observer – 9/25/92 

Ruptured tubing in the propulsion system caused Mars Observer 

to spin out of control. 

Mars 96 – 11/16/96 

A failure during the second ignition of the Proton rocket’s rocket s upper 

stage sent the spacecraft down into the Pacific Ocean.

Nozomi – 7/4/98 

Might be in orbit around the Sun (could be crashing into Spirit any 

minute). 

Mars Climate Orbiter – 12/11/98 

“The The failed translation of English units into metric units in a 

segment of ground-based, ground based, navigation-related navigation related mission software.” 

software. 

Mars Polar Lander/Deep Space 2 – 1/3/99 

“…the “…the 

most probable cause of the failure (of MPL) was the 

generation of spurious signals when the lander legs were deployed deployed 

during descent. The spurious signals gave a false indication that that 

the spacecraft had landed, resulting in a premature shutdown of 

the engines and the destruction of the lander when it crashed on 

Mars.” 

Mars.

Thermal Control Components 

Passive Control: 

• Surface Finishes 

• Insulation 

• Radiators 

Active Control: 

• Heaters & Thermostats 

• Thermal Louvers 

• Heat Pipes

Thermal Control Components: 

Surface Finishes 

Select radiative properties of spacecraft surface to 

achieve an energy balance: 

Qenvironment environment + ∑Qinternal internal = Qre-radiated re radiated 

Two primary surface properties: 

- solar absorptivity, 

absorptivity, 

α 

- IR emissivity, 

emissivity, 

ε 

Surface coatings can be combined for desired average α 

and ε. 

Example: for a sphere located 1 AU from Sun (1367 

W/m 2 ) 

Z93 white: white: 

where α/ε=0.17/0.92, =0.17/0.92, we find T=-90 T= 90°C 

Z306 black: black: 

where α/ε=0.92/0.89, =0.92/0.89, we find T=8°C T=8

Thermal Control Components: Surface Finishes: 

α 

1 

0.9 

0.8 

0.7 

0.6 

0.5 

0.4 

0.3 

0.2 

0.1 

0 

Solar Absorptivity (α) 

Of typical spacecraft surface finish categories 

Optical 

Solar 

Reflectors 

White 

Paints 

Black 

Paints 

Aluminized 

Kapton 


Metallic Anodized Al

Thermal Control Components: Surface Finishes: 

ε 

1 

0.9 

0.8 

0.7 

0.6 

0.5 

0.4 

0.3 

0.2 

0.1 

0 

Optical 

Solar 

Reflectors 

IR Emissivity (ε) 

Of typical spacecraft surface finish categories 

White 

Paints 

Black 

Paints 

Aluminized 

Kapton 


Metallic Anodized Al


Multilayer Insulation (MLI) 

and single layer radiation shields 

Most common type of thermal control elements 

Prevents excessive heat loss or heating by environment or 

other spacecraft components 

Provides radiation protection 

MLI blankets 

• typically protect internal propellant tanks and lines, solid 

rocket motors, cryogenic dewars, dewars, 

& much more 

• not effective in the presence of a gas, must add foam, batt or 

aerogel 

• sensitive to mechanical compression and edge effects and can 

easily decrease from one to two orders of magnitude from its 

ideal value even when the MLI is kept under high vacuum 

condition 

Single layer radiation shields 

• Lesser degree of insulation 

• Lighter 

http://www.nasatech.com/Briefs/Sept03/KSC12092.HTML

Thermal Control Components: Insulation 

Typical MLI Composition 

SMAD III; Fig 11-17A, pg. 437


Radiators 

Most spacecraft waste heat is rejected to space 

by radiators 

Rejects heat by IR radiation from surfaces 

(between 100 and 350 W) 

Heat rejection or heat radiating capability 

increases exponentially with temperature 

Can vary from weighing nothing to 12kg/m 2 

Types of finishes mostly used are quartz mirrors, 

silvered or aluminized teflon & white paint


Heaters & Thermostats 

Heaters: 

Patch heater 

• Resistor between two flexible insulators (Kapton ( Kapton) 

• Redundancy required 

• Can be made into custom shapes 

Cartridge heater 

• Heats blocks of material or high temperature components 

(hydrazine-thruster (hydrazine thruster catalyst beds) 

• Wound resistor enclosed in metallic casing 

SMAD III; Fig 11-19, pg. 440


Thermostats: 

Heaters & Thermostats 

Mechanical thermostat 

• Typically used for a heater that is 

activated only for special events or is on 

all the time 

• Fairly reliable but many used (up to 

several hundred) 

• Large dead band, >4°C >4 

Solid state controller (w/ temp. sensor) 

• Electronic device= higher reliability and 

life expectancy 

• Tight dead bands,


Adjusts exposure of a 

component (radiator) to 

space or to other 

components 

Fully open, it allows the 

rejection of six times as 

much heat if fully closed 

No power required 

Useful when internal 

power dissipation varies 

widely 

Strong conductive path 

between actuator and 

radiator 

Louvers 

“Venetian Venetian-blind blind” type 

SMAD III; Fig 11-21, pg. 442


Heat Pipes 

Wick envelope design transfers heat from one 

end to the other 

Extruded grooves enable this: 

• Trapezoidal 

Easier to make 

• Re-entrant 

Re 

Better capillary 

action 

Trapezoidal 

Easier to make 

entrant 

Better capillary 

SMAD III; Fig 11-22A&B, pg. 445

Application of Thermal Control Components: 

The NASA Rover 

Thermal blankets 

Heat pump and piping capable of 

transporting 150 watts of rover waste 

heat into space 

Radioisotope Heater Units (RHUs ( RHUs) ) 

generate heat through decay of a low- 

grade isotope 

Electrical heaters 

Thermostatically controlled heaters 

Uses a Freon-like Freon like fluid for cooling

Thermal Protection during 

Reentry 

Types of Thermal 

Insulation/Protection: 

1. Reinforced Carbon-Carbon 

2. Ceramic Tiles 

3. Insulation Blankets 

4. Ablative Materials 

http://tps.arc.nasa.gov

Reinforced Carbon-Carbon 

Carbon Carbon 

Used in regions where heat load > 

68W/cm 2 (stagnation regions) and 

temperature

Ceramic Tiles 

http://tps.arc.nasa.gov 

• High Temperature Reusable Surface 

Insulation 

Used in regions where heat load is < 68W/cm 2 

and temperature < 1500K 

Lightweight Ceramic Tile 

Cheaper than RCC 

Difficult to install 

Nomex used as stress isolation 

http://tps.arc.nasa.gov 

• Fibrous Refractory Composite Insulation 

Used in regions where heat load is < 68W/cm 2 

 

and temperature < 900K 

Easier to produce than HRSI 

Denser material than HRSI, comes in two grades 

Difficult to install

Advanced Flexible Reusable 

Surface Insulation Blankets 

Used to replace Ceramic Tiles in regions 

where lower heating occurs 

• Effective up at temperatures < 1050K 

• Easier to produce but has lower 

durability than tiles 

• Decrease in weight and easy to install 

• Also available in two grades 

http://tps.arc.nasa.gov

Ablative Materials 

Used in regions where heat load 

>100W/m 2 

Material chars and flakes or melts off 

• Avcoat 5026-39 5026 39 

• Used in Apollo Missions to shield the Command 

Service Module 

• Dense Material (ρ=1650 ( =1650 kg/m 3 ) 

http://www.spaceaholic.com/avcoat_heatshield_full.jpg

Ablative Materials 

• PICA (Phenolic 

Impregnated Carbon 

Ablator) 

• Used on the Stardust Probe 

(ETA Jan 2006) 

• Lightweight Material (ρ=240 ( =240 

kg/m 3 ) 

http://stardust.jpl.nasa.gov/photo/SC.jpg

Example of Ablation 

Galileo Probe entering Jupiter’s Jupiter s Atmosphere (December 1995) 

http://www.centennialofflight.gov/essay/Evolution_of_Technology/advanced_reentry/Tech20G3.htm

Past Structures 

Best structure is blunt body 

Greater amount of friction 

• Decelerates the most in upper atmosphere 

• Easier to reduce speed to zero for landing 

Causes detached shock 

• Lower skin temperature 

• High temperature is further away from body

Radiation Shielding 

Use of Multi Layer Insulation Blankets to 

cut radiation exposure for payloads 

Effective radiation shield must have high 

Hydrogen atom content, helps 

disperse/absorb radiation 

• Aluminum used in the past as this was 

integrated into structure 

• Reinforced polyethylene being researched by 

NASA (Fiber used to produce bricks for 

radiation shields)

Radiation shield for computers 

The Radiation Shield is a software protection 

system for personal computers in Space running 

Microsoft Windows NT (made by ESA ESTEC) 

The radiation present in Space environment may 

generate parity error in parity protected memory; 

alternatively it can cause system faults or 

application errors. 

The Radiation Shield software is designed to 

handle all these errors for any running application 

thereby allowing it to continue or gracefully 

degrade, without losing significant user data. 

http://www.estec.esa.nl/wmwww/EME/laptops/ntradshld/index.htm

Thermal Material Cost Comparison 

Cost (Dollar/m^2) 

1.40E+05 

1.20E+05 

1.00E+05 

8.00E+04 

6.00E+04 

4.00E+04 

2.00E+04 

0.00E+00 

RCC 

Thermal Material Costs 

AFRSI 

HRSI 

AFRSI LI-900 

2200 

Material 

HRSI 

LI-2200 

Nomex FRSI 12 

RCC AFRSI AFRSI-2200 HRSI LI-900 HRSI LI-2200 Nomex FRSI-12

Thermal Material Density Comparison 

Density (kg/m^3) 

1.80E+03 

1.60E+03 

1.40E+03 

1.20E+03 

1.00E+03 

8.00E+02 

6.00E+02 

4.00E+02 

2.00E+02 

0.00E+00 

RCC 

AFRSI 

Thermal Material Density 

AFRSI 

2200 

HRSI 

LI-900 

HRSI 

LI-2200 

Purchase Cost 

Material 

Nomex 

FRSI 12 

Graphite 

Epoxy 

RCC AFRSI AFRSI-2200 HRSI LI-900 HRSI LI-2200 Nomex FRSI-12 Graphite Epoxy

Temperature (K) 

Thermal Material Max Temperature Limit Comparison 

2.50E+03 

2.00E+03 

1.50E+03 

1.00E+03 

5.00E+02 

0.00E+00 

RCC 

AFRSI 

Thermal Material Max Temperature 

AFRSI 

2200 

HRSI 

LI-900 

Material 

HRSI 

LI-2200 

Nomex 

FRSI 12 

RCC AFRSI AFRSI-2200 HRSI LI-900 HRSI LI-2200 Nomex FRSI-12 Graphite Epoxy 

Graphite 

Epoxy

Thermal Material Installation Time Comparison 

Installation Time (s/m^2) 

4.00E+06 

3.50E+06 

3.00E+06 

2.50E+06 

2.00E+06 

1.50E+06 

1.00E+06 

5.00E+05 

0.00E+00 

Thermal Material Installation Time 

RCC AFRSI AFRSI 

2200 

HRSI 

LI-900 

HRSI 

LI-2200 

Nomex 

FRSI 12 

RCC AFRSI AFRSI-2200 HRSI LI-900 HRSI LI-2200 Nomex FRSI-12

Density (kg/m^3) 

5.00E+03 

4.50E+03 

4.00E+03 

3.50E+03 

3.00E+03 

2.50E+03 

2.00E+03 

1.50E+03 

1.00E+03 

5.00E+02 

0.00E+00 

Structural Material Density Comparison 

Aluminum 

2024 

Structural Material Density 

Aluminum 

2219 

Material 

Titanium 

Al 2024-T8XX 

AL 2219-T8XX 

Titanium

Tensile Strength (Pa) 

Structural Material Tensile Strength Comparison 

1.20E+09 

1.00E+09 

8.00E+08 

6.00E+08 

4.00E+08 

2.00E+08 

0.00E+00 

Aluminum 

2024 

Structural Material Tensile Strength 

Aluminum 

2219 

Material 

Al 2024-T8XX Al 2219-T8XX Titanium Duplex Annealed 

Titanium

Comprehensive Thermal Material Table 

Units RCC HRSI LI-900 HRSI LI-2200 Normex FRSI-12 AFRSI AFRSI-2200Epoxy 

Purchase Cost $/m^2 1.29E+05 1.23E+04 1.25E+04 1.72E+03 1.25E+04 3.55E+03 1.25E+04 

Installation Time s/m^2 3.72E+05 3.53E+06 3.53E+06 2.13E+04 3.53E+06 2.36E+05 3.53E+06 

Inspection/Repair Time per Flight s/m^2 5.43E+03 8.14E+04 8.14E+04 3.49E+03 8.14E+04 3.72E+04 8.14E+04 

Replacement Fraction per Flight - 1.30E-03 2.50E-03 2.50E-03 2.80E-02 2.50E-03 1.80E-02 2.50E-03 

Reuse Flight Limit (# of Flights) - 3.30E-01 1.00E+02 1.00E+02 0.00E+00 1.00E+02 0.00E+00 1.00E+02 

Density kg/m^3 1.58E+03 1.44E+02 3.52E+02 8.65E+01 1.92E+02 9.61E+01 3.52E+02 1.58E+03 

Thermal Conductivity (Thru-the-thickness) W/m-K 4 4.76E-02 7.44E-02 4.06E-02 5.30E-02 3.29E-02 7.44E-02 0.6 

Thermal Conductivity (In plane) W/m-K 6 6.75E-02 0.1 7.96E-02 0.1 3 

Specific Heat J/kg-K 7.12E+02 6.28E+02 6.28E+02 1.31E+03 7.12E+02 7.41E+02 6.28E+02 8.33E+02 

Single Use Temperature Limit K 2.03E+03 1.76E+03 1.81E+03 6.44E+02 1.81E+03 1.09E+03 1.81E+03 5.89E+02 

Emissivity 0.78 8.80E-01 9.00E-01 8.00E-01 9.20E-01 8.70E-01 9.00E-01 

Multiple Use Temprature Limit K 1.92E+03 1.59E-03 1.64E+03 5.06E+02 1.64E+03 9.22E+02 1.64E+03 

Tensile Strength (Thru-the-thickness) Pa 1.65E+05 5.03E+05 1.58E+05 5.03E+05 

Tensile Strength (In-Plane) Pa 4.69E+05 1.25E+06 1.77E+06 1.25E+06 

Tensile Modulus (Thru-the-thickness) Pa 4.83E+07 1.86E+08 6.89E+07 1.86E+08 

Tensile Modulus (In Plane) Pa 1.72E+08 5.52E+08 3.45E+08 5.52E+08 

Compressive Strength (Thru-the-thickness) Pa 1.93E+05 8.96E+05 9.11E+05 8.96E+05 

Compressive Strength (In-Plane) Pa 4.83E+05 1.59E+06 1.83E+06 1.59E+06 

Coefficient of Thermal Expansion (In-Plane) 1/K 1.31E-06 4.05E-07 4.83E-07 1.30E-06 4.83E-07 

Dielectric Constant - 1.13E+00 1.30E+00 1.20E+00 1.30E+00 

Loss Tangent - 4.00E-04 1.60E-03 9.00E-04 1.60E-03

Comprehensive Structural Material Table 

Units Alum. 2024 Alum. 2219 Titanium 

Purchase Cost $/m^2 

Installation Time s/m^2 

Inspection/Repair Time per Flight s/m^2 

Replacement Fraction per Flight - 

Reuse Flight Limit (# of Flights) - 

Density kg/m^3 2.80E+03 2.81E+03 4.50E+03 

Thermal Conductivity (Thru-the-thickness) W/m-K 1.45E+02 1.19E+02 2.19E+01 

Thermal Conductivity (In plane) W/m-K 

Specific Heat J/kg-K 8.16E+02 8.62E+02 5.22E+02 

Single Use Temperature Limit 

Emissivity 

K 4.50E+02 4.50E+02 1.95E+03 

Multiple Use Temprature Limit K 

Tensile Strength (Thru-the-thickness) Pa 4.55E+08 4.55E+08 1.11E+09 

Tensile Strength (In-Plane) Pa 4.00E+08 3.50E+08 

Tensile Modulus (Thru-the-thickness) Pa 

Tensile Modulus (In Plane) Pa 

Compressive Strength (Thru-the-thickness) Pa 

Compressive Strength (In-Plane) Pa 

Coefficient of Thermal Expansion (In-Plane) 1/K 

Dielectric Constant - 

Loss Tangent -

Airbags 

http://cmex.arc.nasa.gov/CMEX/data/catalog/MarsSurveyorRover2003/ParachuteAirbagsLanding.html 

Vectran M Vectran HS 

LCP Fiber LCP Fiber 

Density kg/m3 1400 1400 

Moisture Absorption at Equilibirum Max (%) 0.1 0.1 

Tensile Strength Ultimate Mpa 1110 3025 

Elongation at Break % 2 3.5 

Tensile Modulus GPa 52.4 68.6 

Dielectric Constant 3.3 3.3 

Melting Point K 549 603

Parachutes 

http:// marsrovers.jpl.nasa.gov/mission/parachute2.html 

Made of polyester and nylon 

Bridle is made out of Kevlar 

In case of Exploration Rover, the parachute 

resisted approximately 80,100~84,600 N of 

force; parachutes were inflated at Mach 6 and 

decreased the velocity to Mach 2 

Size and materials depend on atmospheric 

density, velocity, parachute drag area and 

mass

Conclusion 

Structural design depends on weight 

and structural stiffness over a wide 

range of temperatures 

Learn from the past; Select best 

materials to accommodate needs 

Major factors in choosing materials 

Cost, Installation Time, Working 

Temperature, Strength

http://www.mse.berkeley.edu/classes/matsci102/F01reports/tiles.pdf 

http://www.mse.berkeley.edu/classes/matsci102/F01reports/tiles.pdf 

http://spaceflight.nasa.gov/shuttle/reference/shutref/orbiter/tps/hrcitiles.html 

http://spaceflight.nasa.gov/shuttle/reference/shutref/orbiter/tps/hrcitiles.html 

http://www-pao.ksc.nasa.gov/kscpao/nasafact/tps.htm 

http://www pao.ksc.nasa.gov/kscpao/nasafact/tps.htm 

http://er6s1.eng.ohio-state.edu/mse/mse205/lectures/chapter20/chap20.pdf 

http://er6s1.eng.ohio state.edu/mse/mse205/lectures/chapter20/chap20.pdf 

http://www.eng.iastate.edu/explorer/topics/spacecoat/HRSI.htm 

http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/stsref 

http://science.ksc.nasa.gov/shuttle/technology/sts newsref/stsref-toc.html#sts_ov 

toc.html#sts_ov 

http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/sts_sys.html 

http://science.ksc.nasa.gov/shuttle/technology/sts newsref/sts_sys.html 

http://www.floridatoday.com/columbia/columbiastory2A45646A.htm 

http://www.kc4cop.bizland.com/shuttle_tiles_1.htm 

http://www.tc.cornell.edu/Research/CMI/RLVsource/tps.html 

http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/sts_coord.html#sts 

http://science.ksc.nasa.gov/shuttle/technology/sts newsref/sts_coord.html#sts-wing wing 

http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/sts 

http://science.ksc.nasa.gov/shuttle/technology/sts newsref/sts-rcs.html#sts 

rcs.html#sts-rcs rcs-heaters heaters 

http://marsrovers.jpl.nasa.gov/mission/spacecraft_edl_aeroshell.html 

http://marsrovers.jpl.nasa.gov/mission/spacecraft_edl_aeroshell. html 

http://www.matweb.com/search/SpecificMaterial.asp?bassnum=MA2240 

http://www.matweb.com/search/SpecificMaterialPrint.asp?bassnum=CREADE021 

http://www.matweb.com/search/SpecificMaterialPrint.asp?bassnum=CREADE021 

http://www.matweb.com/search/SpecificMaterialPrint.asp?bassnum=PCELAN00 




http://www.matweb.com/search/SpecificMaterialPrint.asp?bassnum=PHEH32 

http://www.matweb.com/search/SpecificMaterialPrint.asp?bassnum=PHEH32 

http://tpsx.arc.nasa.gov 

http://www.cnn.com/interactive/space/0308/earth.mars/frameset.exclude.html 

http://www.cnn.com/interactive/space/0308/earth.mars/frameset.exclude.html 

http://www.popsci.com/popsci/aviation/article/0,12543,0459759-1,00.htmlcmex 

http://www.popsci.com/popsci/aviation/article/0,12543,0459759 1,00.htmlcmex-www.arc.asa.gov/CMEX/index.htmlmars.tv/missions.html 

www.arc.asa.gov/CMEX/index.htmlmars.tv/missions.html 

http://www.floridatoday.com/news/space/stories/2003b/121403marsside.htm 

http://www.floridatoday.com/news/space/stories/2003b/121403marsside.htm 

http://spaceflightnow.com/mars/mera/031231mer.html 

http://www.polaris.iastate.edu/EveningStar/Unit7/unit7_sub2.htm 

http://www.nasa.gov/columbia/caib/PDFS/VOL1/PART01.PDF 

http://calspace.ucsd.edu/Mars99/docs/library/mars_exploration/robotic_missions/orbiters/older_missions15.html 

http://calspace.ucsd.edu/Mars99/docs/library/mars_exploration/robotic_missions/orbiters/older_missions15.html 

http://athena.cornell.edu/mars_facts/past_missions_70s.html 

http://www.bbc.co.uk/science/space/exploration/missiontimeline/mars2.shtml 

http://www.bbc.co.uk/science/space/exploration/missiontimeline/mars2.shtml 



http://www.solarviews.com/eng/mars.htm 

http:// www.solarviews.com/eng/mars.htm 

http://www.marssociety.at/ballon5.html 

http://www.luke.com/marsgeo/volcanic5.html 

http://abenteuer-universum.vol4u.de/mars.html 

http://abenteuer universum.vol4u.de/mars.html 

Text: mars surface ionizing radiation environment: need for validation. validation. 

J.W. Wilson, etc 

http://vojager.cet.edu/iss/techcheck3/radiation.html 

Text: materials for shielding astronauts from the hazards of space space 

radiation. J.W. Wilson etc. 

Textbook: Space Mission Analysis and Design 

http://liftoff.msfc.nasa.gov/academy/space/atmosphere.html 

http://www.madsci.org/posts/archives/dec97/880000587.As.r.html 


http://marsrovers.jpl.nasa.gov/mission/spacecraft_edl_parachute.html 

http://marsrovers.jpl.nasa.gov/mission/spacecraft_edl_parachute. html 

http://marsrovers.jpl.nasa.gov/mission/spacecraft_edl_airbags.html 

http://marsrovers.jpl.nasa.gov/mission/spacecraft_edl_airbags.html 

Text: NASA Facts Orbiter Thermal Protection System 

http://www.e-composites.com/commonFiles/engineering_material_properties.asp 

http://www.e composites.com/commonFiles/engineering_material_properties.asp 

http://www.corrosion-doctors/matselect/cost.htm 

http://www.corrosion doctors/matselect/cost.htm 

http://www.physics.purdue.edu/cmsfpixs/mechanics/panel_support/materials/aluminum 

http://www.physics.purdue.edu/cmsfpixs/mechanics/panel_support/materials/aluminum-alloy 

alloy 

http://www.lowden-metals.co.uk/funnace/cfrc.html 

http://www.lowden metals.co.uk/funnace/cfrc.html 

http://nova.stanford.edu/projects/mod/late.html 

http:// nova.stanford.edu/projects/mod/late.html 

http://slate.msn.com/id/2093779/ 

http://www.badastronomy.com/phpBB/viewtopic.php?t 

http:// www.badastronomy.com/phpBB/viewtopic.php?t=9535 =9535 

http://eto.msfc.nasa.gov/challenge.html#tps 

http://lasp.colorado.edu/snoe/lib/sc_struct.html 

http://www.aa.washington.edu/research/isru/ares/ares.htm 

http:// www.aa.washington.edu/research/isru/ares/ares.htm 

http://www.centennialofflight.gov 

http:// www.centennialofflight.gov/ / 

http://esapub.esrin.esa.it/bulletin/bullet87/paroli87.htm 

http://www.matweb.com 


http://www.nasatech.com/Briefs/Sept03/KSC12092.HTML 

http://www.estec.esa.nl/wmwww/EME/laptops/ntradshld/index.htm 

Phil Herlth, a NASA employee 

References

Thank You

Structures / Thermal Group

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