Thermal protection materials and systems (TPS) have been critical to fulfilling humankinds desire to explore space. Composite and ceramic materials have enabled the early missions to orbit, the moon, the space station, Mars with robots, and sample return. Crewed missions to Mars are being considered, and this places even more demands on TPS materials. This talk will give some history on the materials used for earth and planetary entry and the demands placed upon such materials. TPS needs for future missions, especially to Mars, will be identified and potential solutions discussed

Johnson, Sylvia M.

NASA Technical Reports Server

Thermal Protection Materials and 
Systems: Past and Future
Sylvia M. Johnson
NASA Ames Research Center
40th International Conference and Exposition 
on Advanced Ceramics and Composites
Daytona Beach
January 25, 2015
https://ntrs.nasa.gov/search.jsp?R=20160001151 2019-08-31T04:20:37+00:00Z
Apollo Heatshield: After Entry!
Material: AVCOAT
Atmospheric entry  is tough on materials! 2
Outline
• Introduction to Thermal Protection Systems (TPS)
• Reusable TPS
- Shuttle materials
- Current reusable material
• UHTCs 
• Ablative TPS
- Recent materials
- Materials selection
- Orion TPS
• Challenges for the future
New materials/concepts
3
Introduction to TPS  for NASA
• NASA uses TPS for entry vehicles that carry people, 
cargo, experiments, samples and/or instruments
- Entries usually involve descending into the atmosphere followed by  
a trajectory that aims to burn off energy and result in a controlled 
landing.
- Not usually ballistic entries
- Not long hypersonic operations in atmosphere.
- Entry to Earth, Mars, Venus, gas and ice giant planets
4
TPS Development
• NASA Ames focuses on:
- Qualifying and certifying TPS for current missions 
- Developing new TPS for upcoming missions
• Approaches to TPS development differ with risk —
crewed vs. robotic missions:
- Crewed
 Loss of life must be avoided
 What must be done to qualify and certify TPS?
- Robotic missions
 Can take more risk
 But scientific knowledge can be lost too
• Goal for all TPS is efficient and reliable performance
• Need to understand materials to enable design and use
5
Entry Heating Parameters
• Reentry heating : 2 primary sources
- Convective heating from both the flow of hot gas past the surface of the 
vehicle and catalytic chemical recombination reactions at the surface
- Radiation heating from the energetic shock layer in front of the vehicle
• Heating depends on reentry speed (V), vehicle effective radius (R), and 
atmospheric density (ρ) 
• As reentry speed increases, both convective and 
radiation heating increase
- Radiation heating dominates at high speeds
• As vehicle radius increases, 
convective heating decreases,
but radiation heating increases
V
R
5.0
3







R
Vqconv


Convective Heating
5.02.18 RVqrad 
Shock Radiation Heating
6
Thermal Protection Systems
• Protect vehicle structure and contents (people and things) from the heat of 
entry through an atmosphere
• Rely on material’s response to environment
• Response depends on
- Material properties
- Configuration of the system
- Specific conditions (heat flux, pressure, flow)
• Physical Forms: rigid, conformable, flexible
One size does not fit all!
Different TPS for different vehicles, location on vehicles, 
and mission conditions
Goal of all TPS is reliable and 
efficient performance
Specifically addresses challenges of mass reduction and 
reliability
7
Rigid, Conformable and Flexible TPS
• Rigid – fabricated in a rigid form and usually 
applied in a tiled configuration to a rigid 
substructure
• Conformable – fabricated in a flexible form and 
shaped to a rigid substructure; final form may 
be rigid or compliant
• Flexible – fabricated and used in a flexible form, 
where flexibility is an essential component of 
the heatshield, e.g., deployable systems, 
stowable systems
• Woven – can be any of the above
Physical Forms of TPS
8
Reusable vs. Ablative TPS
9
Energy management through storage and re-radiation — material unchanged
When exposed to atmospheric 
entry heating conditions, 
surface material will heat up 
and reject heat in the following 
ways:
• Re-radiation from the 
surface and internal storage 
during high heating 
condition
• Re-radiation and convective 
cooling under post-flight 
conditions
Insulative/Reusable TPS
radiation 
flux out
convective 
flux
radiation 
flux in
boundary layer 
or shock layer
high emissivity 
coating
low conductivity 
insulation TPS
backup or 
structure material
free stream
conduction 
flux
10
Ames-Developed Thermal Protection Materials 
Adopted on Shuttle
AIM-22 Tile
RCG Coating
AFRSI Blanket
FRCI-12 Tile
TUFI/AETB Tile
Gap Fillers
11
Reusable TPS: Tiles
• Effort started in 1970’s by ARC to provide NASA with TPS materials and 
processing expertise
• Insulation materials used to protect the aluminum sub-structure of the 
shuttle.
• High purity silica, aluminoborosilicate, and alumina fibers
• LI-900, FRCI-12, AETB-8
• Open porous structure
• Used on over 100 shuttle missions
100 mm
AETB (35% Al2O3) Tile
10mm
10mm
10mm
Silica fibers
Alumina fibers
Nextel® fibers
Starting materials for tiles
Tiles are  heterogeneous with regions of low density 
and clumps of fibers with some non-fibrous inclusions 12
Reusable TPS: Tiles and Coatings
400 mm
400 mm
RCG Coating TUFI Coating
• RCG (Reaction Cured Glass) is a thin 
dense high emittance glass coating on the 
surface of shuttle tiles
• Poor impact resistance
• TUFI (Toughened Unipiece Fibrous 
insulation) coatings penetrate into the sample
• Porous but much more impact resistant 
system
• Silica-based fibers
• Mostly empty space-
>90%porosity
100 mm
Density: 0.14 to 0.19 g/cm3 “Space Shuttle Tile”
13
Shuttle Flight Testing of TUFI Tiles in Base 
Heatshield
TUFI/AETB-8 Tiles Undamaged After 
Three Flights
Silica-based Tile
TUFI tiles used on base heatshield of 
Shuttle to protect against damage 
from debris incurred during liftoff
RCG Hybrid
Overcoat
Impregnated surface treatment
14
C
a
p
B
a
s
e
 I
n
s
u
la
to
r ROCCI
Fibrous Insulation 
Graded Surface Treatment
Schematic of TUFROC TPS
TUFROC TPS
(Toughened Unipiece Fibrous Reusable Oxidation Resistant Ceramic)
• Developed TUFROC for X-37 application
• Advanced TUFROC developed recently
• Transferred technology to Boeing and others
• System parameters:
- Lightweight (similar to LI-2200) 
- Dimensionally stable at surface temperatures up to1922 K
- High total hemispherical emittance (0.9)
- Low catalytic efficiency
- In-depth thermal response is similar to single piece Shuttle-type fibrous insulation    
X-37  Reentry Vehicle
Wing leading edgeNose cap
Control surface 
15
ROCCI Carbonaceous  Cap 
- Silicon-oxycarbide phase slows oxidation
- HETC, treatment near surface slows 
oxidation and keeps emissivity high (ε ~ 0.9)
- Coated with borosilicate reaction cured 
glass (    RCG    ) for oxidation resistance
AETB Silica Insulating Base
- Solved thermo-structural issues by adding boron 
oxide (B2O3) and alumino-borosilicate fibers, 
which also improved mechanical strength
- Increased temp capability to 2500+ °F by  
adding alumina (Al2O3) fiber
Standard TUFROC
2 Piece Approach
Re-radiate enough heat so that conduction through 
- Cap is within temp limits of the insulating Base 
- Base is within temp limits of the Vehicle
AETB Insulating Base
re-radiation∝ εT4
significantly reduces heat  
conducted to the vehicle
max temp: 2600 °F
3000
2500
400
200
R E - E N T R Y
H E A T I N G
Max
Temp 
(°F)
VEHICLE STRUCTURE
heat conduction
ROCCI Cap
maintains outer mold line
max temp: 3000 °F
TUFROC
16
Flight Proven Standard TUFROC
TUFROC spans USAF X-37b wing leading edge
- NASA developed Standard TUFROC and  
transferred it to X-37b Prime - Boeing
- Enabling technology for critical USAF Program
- 3 successful missions, 4th mission in progress
Reusability of Standard TUFROC? ⇒ Advanced
X-37b Preparing for 1st launch, Apr 2010
12/8/2010
X-37b after 224 days (90 million miles) in orbit, Dec 2010
TUFROCTUFROC
TUFROC on X-37
Sharp Leading Edge Energy Balance
• Insulators and UHTCs manage energy in different 
ways:
- Insulators store energy until it can be eliminated in the same 
way as it entered
- UHTCs conduct energy through the material and reradiate it 
through cooler surfaces
Dean Kontinos, Ken Gee and Dinesh Prabhu. “Temperature Constraints at the Sharp Leading Edge of a 
Crew Transfer Vehicle.” AIAA 2001-2886 35th AIAA Thermophysics Conference, 11-14 June 2001, 
Anaheim CA                    
UHTC
High 
Thermal 
Conductivity
Sharp Nose
Sharp Nose
Leading Edges
18
Ultra High Temperature Ceramics (UHTCs) : A 
Family of Materials
• Borides, carbides and nitrides of 
transition elements such as hafnium, 
zirconium, tantalum and titanium
• Some of highest known melting 
points
• High hardness, good wear 
resistance, good mechanical strength
• Good chemical and thermal stability 
under certain conditions
- High thermal conductivity
- Good thermal shock resistance
• The microstructure of UHTCs clearly 
shows their composite nature
- Distribution of material phases
- Flaw size and distribution
Teledyne Scientific
19
Ultra High Temperature Ceramics (UHTCs) : A 
Family of Materials
• Borides, carbides and nitrides of 
transition elements such as hafnium, 
zirconium, tantalum and titanium
• Some of highest known melting 
points
• High hardness, good wear 
resistance, good mechanical strength
• Good chemical and thermal stability 
under certain conditions
- High thermal conductivity
- Good thermal shock resistance
• Considerable effort in many 
institutions to improve properties and 
processing of UHTCs
Hf-B Phase Diagram
20
Energy management through material consumption
When exposed to 
atmospheric entry heating 
conditions, material will 
pyrolyze (char), and reject 
heat in the following ways:
•Endothermic decomposition 
of polymer
•Blowing of ablation 
products into the boundary 
layer reduces convective 
heating
•Formation of char layer and 
re-radiation
Ablative TPS
21
Phenolic Impregnated Carbon Ablator (PICA)
Drying  Cycle
Resin Impregnation
Fiberform™ before impregnation
PICA: Fiberform™
with phenolic resin
Phenolic Resin
PICA Arc Jet 
Model
Carbon Fiberform™
PICA has low density (~0.27g/cm3) 
and is an efficient ablator at high 
heat fluxes
Processing Detail
22
PICA Applications
PICA was the enabling TPS 
material for the Stardust 
mission where it was used as 
a single piece heatshield
Stardust sample return capsule post flight with PICA 
as the forebody TPS. (0.8m diameter) MSL Heat Shield (4.5m diameter)
PICA was the primary heatshield 
for Mars Science Lab (MSL) and 
a variant is used in SpaceX’s
Dragon cargo vehicle in a tiled 
configuration
23
Selection of Appropriate Material
• Historical approach:
- Use heritage materials: “It’s worked before…”
- Risk-reduction strategy
- Limited number of flight-qualified ablative materials
- Different vehicle configurations and reentry conditions (need to 
qualify materials in relevant environments)
• As missions become more demanding, we need higher 
capability materials — necessary to have a robust 
research and development program
• Reusable and ablative materials are both needed
• Must test materials in relevant environments
• Provide path for insertion/use of new materials
24
Need for Arc Jet Testing
• Arc jet testing is the best ground-
based method of evaluating a 
material’s oxidation/ablation response 
in re-entry environments
• Oxidation/ablation behavior on heating 
in static or flowing air at ambient 
pressures is likely to be significantly 
different than in a re-entry 
environment.
- O2 and N2 may be dissociated
 Catalycity of the material
 Recombination of O and N atoms 
adds to surface heating
- Stagnation pressures may be <1 atm.
25
Orion 
• The Orion spacecraft will take astronauts beyond low Earth orbit 
(LEO) to deep space.
- emergency abort capability,
- sustain the crew
- provide safe re-entry from deep space.
• Exploration Flight Test-1, an uncrewed mission flew in 2014.
- Orion travel farther into space than any spacecraft had gone in more than 40 years.
- EFT-1 data used to influence design decisions, validate existing computer models and 
innovative new approaches to space systems development, as well as reduce overall 
mission risks and costs.
• What TPS was used?
- Decision was made to use Avcoat, material first used on Apollo 
http://www.nasa.gov/images/content/545955main_mp
cvstack_full.jpg
26
Orion TPS: AVCOAT
• Avcoat was used on the Apollo 
vehicles: “heritage” material
• Consists of a honeycomb filled with 
an ablative mixture
• Complex material requiring hand 
assembly
• Next Orion flight will use a variant 
of Avcoat without the honeycomb
Avcoat arc jet models: pre 
and post test
27
Destinations and Challenges
• Saturn and Venus: robotic missions
- Very extreme environments, especially Venus atmosphere
- Saturn: very large, very high heating on entry
• Mars: robotic and crewed mission
- Crew requires large amount of cargo
- Crew to and from surface separately
28
TPS Selection 
• Entry into outer planets/ Venus
- Large aeroshells for deceleration
• Entry into Mars
- Sky crane approach of MSL/Curiosity 
not feasible for loads>1.5mt to Mars
- Balloons / parachutes not very effective
- Need large aeroshell
• High speed entry into Earth’s 
atmosphere
- Direct trip/ entry: entry speed> 13.5km/s 
- Orion vehicle: need more capable TPS
- Inspiration Mars proposed very small 
reentry vehicle: lower heat flux, current 
TPS
• Scenarios have differing degrees of 
risk to humans—length of time in 
space, entry speeds, g forces, 
hazard of changing vehicles
Planet
Mission
Studies
Peak Heat 
Flux 
Range
(W/cm2)
Pressure 
Range 
(atm)
Heat Load 
Range
(kJ/cm2)
Venus1 2400 -
4900 
4 - 9 11 - 12
Saturn2 1900 -
7700
2 - 9 80 - 272
http://www.nasa.gov/exploration/systems/orion/gallery/index.html?id=341169
1. Prabhu, D.K., et. al.; IEEE Aerospace Conference, Big Sky, MT, March 2-9, 2013
2. Allen, G. A. and Prabhu, D. K.; private communication
29
Future Missions: TPS Availability and “Gap”
• Carbon phenolic
- high heat fluxes 
• PICA/other ablators
- < ~ 2,000 W/cm2
No efficient TPS for the gap from ~1000W/cm2 to 10,000W/cm2
Historical TPS Mass Fraction 
by Heat Flux and Pressure
Outer planet and 
sample return 
missions
SLA
PICA
Carbon Phenolic
30
3D Woven TPS
An approach to the design and manufacturing of ablative TPS by the 
combination of weaving  precise placement of fibers in an optimized 
3D woven manner and then resin transfer molding when needed 
• Design TPS for a specific mission 
• Tailor material composition by weaving together different 
types of fibers and by exact placement using computer 
controlled, automated, 3-D weaving technology
• One-step process for making a mid-density dry woven 
TPS
• Ability to infiltrate woven preforms with polymeric resins 
for highest density TPS to meet more demanding thermal 
requirements
Blended Yarn Resin infused 31
0.00	
0.50	
1.00	
1.50	
2.00	
2.50	
3.00	
3.50	
4.00	
-35	 -30	 -25	 -20	 -15	 -10	 -5	 0	
A
re
al
	M
as
s,
	g
/c
m
2
	
Entry	Angle,	degree	
0.00	
0.50	
1.00	
1.50	
2.00	
2.50	
3.00	
3.50	
4.00	
-35	 -30	 -25	 -20	 -15	 -10	 -5	 0	
A
re
al
	M
as
s,
	g
/c
m
2
	
Entry	Angle,	o	
HEEET:	V=10.8	km/s	
HEEET:	V=11.6	km/s	
Carbon	Phenolic:	V=10.8	km/s	
Carbon	Phenolic:	V=11.6	km/s	
0.00	
0.50	
1.00	
1.50	
2.00	
2.50	
3.00	
3.50	
4.00	
-35	 -30	 -25	 -20	 -15	 -10	 -5	 0	
A
re
al
	M
as
s,
	g
/c
m
2
	
Entry	Angle,	o	
HEEET:	V=10.8	km/s	
HEEET:	V=11.6	km/s	
Carbon	Phenolic:	V=10.8	km/s	
Carbon	Phenolic:	V=11.6	km/s	
Acreage Gap 
filler
Potential Mass Savings!
• Improved mass efficiency of woven TPS material for Venus entry 
- More mass for instrumentation
- Lower G loads
32
Deployable Heat Shield Concept 
Carbon Fabric Aeroshell
Rigid Nose
Deployment system
TPS:
• 6 layers of carbon fiber weave (3D weave)
• Has to withstand aerodynamic and 
aerothermal loads.
• Medium Heat Rate Capability (250 W/cm2)
Test model of deployable system
Current concepts for Venus 
exploration
Potential for expansion to Mars entry 
(~16m diameter)
Large sizes will place significant 
demands on structure and 
mechanisms
33
Inflatable Heatshield Concept
TPS
Outer heat resistant layer
Middle insulating layer
Inner non-permeable layer
Example: SiC/carbon felt 
layers impregnated with 
pyrogel
Inflatable aeroshell covered in TPS
High Energy Atmospheric Reentry Test (HEART) 
Current testing in few meter diameter range.
16-20m required for Mars entry and ~20t load.
Issue: long term leakage.  
Inflatable structure  with a flexible TPS
34
Materials for High Speed Earth Entry
• Capability is related to the density of the materials
- Using low density materials at very high heat fluxes leads to rapid recession
• PICA/advanced PICA—capable up to 11km/sec (lunar return), probably more 
capable but not fully tested
- Stardust (<1m) came in at 12.6m/s (1200W/cm2)
- Testing up to 2000W/cm2 in progress
• Woven TPS: tested up to 8000W/cm2
• Conformable PICA capable maybe up to 1000W/cm2, but recession/shear need to 
be better characterized 
• Need new materials and concepts that allow for
- Tailored materials—different properties through thickness to reduce mass and 
improve performance
- Handle radiative entry heating as well as convective heating:surface treatments to 
reflect radiation
- Anti-catalytic coatings that prevents release of heat at the surface
Key is to balance design and materials to make efficient and 
reliable TPS for space exploration. 35
TPS Solutions Availability
• Potentially available for Venus and Saturn (3D woven TPS, 
deployable aeroshells)
• Potentially available for landing cargo on Mars
- Deployable or inflatable concepts for heavy loads
• Current materials probably satisfactory for landing small 
human craft on Mars 
- PICA, existing ablators
- Could also use deployable or inflatable concepts
• Returning people or samples to earth 
- Current materials not sufficient for high speed (>~13.5km/s) 
entry of Orion type vehicle
- Can design mission to involve transfer or use of smaller 
vehicles but involves risk and complexity
36
Summary
• Select thermal protection material based on entry 
conditions
- Ablative materials were used in early vehicles
- Insulative materials were used for Shuttle as it was a reusable 
system
- Ablative materials used in current systems for heatshield
- Insulators still used on certain vehicles (X37b) an on back shell of 
capsules such as Orion where heat flux is lower.
• Outer planets and high speed earth entry require more 
capable materials 
- 3d woven materials for heat shields
- Flexible and deployable materials
• On-going development of materials for NASA missions
37
Remaining TPS Needs
• Improve existing concepts and develop 
higher capability materials
• Capability to test in relevant environments 
and provide data for modeling
• Characterize materials to understand 
behavior
• Develop computational materials 
approaches to design, processing and 
lifetime prediction
38
Acknowledgements
• Thomas Squire (NASA-ARC)
• Joseph Conley (NASA-ARC)
• Mairead Stackpoole (NASA-ARC)
• Alan Cassell (NASA-ARC)
• All the people who have worked on these technologies 
over many years
39
National Aeronautics and 
Space Administration
Ames Research Center
Entry Systems and Technology Division
Optimized LI-900/TUFI
System Schematic
LI-900 Tile
Toughened Surface Treatment
RCG Hybrid
Overcoat
This system reduces the weight of TUFI/LI-900 to an acceptable level by 
limiting the area where the surface treatment is applied while retaining the 
improved damage resistance of the TUFI system.
41
X-37B after Landing
TUROC is on Leading Edges
42


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