SIST-TP CEN/CLC/TR 17603-31-16:2021

SLOVENSKI STANDARD
01-oktober-2021
Vesoljska tehnika - Priročnik o toplotni zasnovi - 16. del: Sistem toplotne zaščite
Space Engineering - Thermal design handbook - Part 16: Thermal Protection System
Raumfahrttechnik - Handbuch für thermisches Design - Teil 16: Wärmeschutzsystem
Ingénierie spatiale - Manuel de conception thermique - Partie 16: Protection Thermique
des véhicules spatiaux
Ta slovenski standard je istoveten z: CEN/CLC/TR 17603-31-16:2021
ICS:
49.140 Vesoljski sistemi in operacije Space systems and
operations
2003-01.Slovenski inštitut za standardizacijo. Razmnoževanje celote ali delov tega standarda ni dovoljeno.

TECHNICAL REPORT
CEN/CLC/TR 17603-31-
RAPPORT TECHNIQUE
TECHNISCHER BERICHT
August 2021
ICS 49.140
English version
Space Engineering - Thermal design handbook - Part 16:
Thermal Protection System
Ingénierie spatiale - Manuel de conception thermique - Raumfahrttechnik - Handbuch für thermisches Design -
Partie 16 : Protection Thermique des véhicules Teil 16: Thermalschutzsysteme
spatiaux
This Technical Report was approved by CEN on 28 June 2021. It has been drawn up by the Technical Committee CEN/CLC/JTC 5.

CEN and CENELEC members are the national standards bodies and national electrotechnical committees of Austria, Belgium,
Bulgaria, Croatia, Cyprus, Czech Republic, Denmark, Estonia, Finland, France, Germany, Greece, Hungary, Iceland, Ireland, Italy,
Latvia, Lithuania, Luxembourg, Malta, Netherlands, Norway, Poland, Portugal, Republic of North Macedonia, Romania, Serbia,
Slovakia, Slovenia, Spain, Sweden, Switzerland, Turkey and United Kingdom.

CEN-CENELEC Management Centre:
Rue de la Science 23, B-1040 Brussels
© 2021 CEN/CENELEC All rights of exploitation in any form and by any means Ref. No. CEN/CLC/TR 17603-31-16:2021 E
reserved worldwide for CEN national Members and for
CENELEC Members.
Table of contents
European Foreword . 5
1 Scope . 6
2 References . 7
3 Terms, definitions and symbols . 8
3.1 Terms and definitions . 8
3.2 Abbreviated terms. 8
4 Introduction . 9
4.1 General .9
4.2 Classification of thermal protection systems . 10
5 Ablative systems . 14
5.1 General . 14
5.2 Ablative materials . 14
5.3 Basic analysis . 15
5.3.1 Surface equilibrium . 16
5.4 Existing systems . 19
5.4.1 Galileo probe . 19
6 Radiative systems . 23
6.1 General . 23
6.2 Radiative materials . 23
6.3 Existing systems . 24
6.3.1 Space shuttle . 24
6.4 Other developments . 35
6.4.1 X-38 . 35
Bibliography . 54

Figures
Figure 4-1: Velocity-altitude map for the Space Shuttle. Lifting re-entry from orbit. . 9

Figure 4-2: Summary of re-entry trajectories. From East (1991) [6]. . 10
Figure 4-3: Sketch of an ablative thermal protection system. . 11
Figure 4-4: Sketch of a radiative thermal protection system. 11
Figure 4-5: Sketch of a transpiration thermal protection system. . 12
Figure 4-6: Typical transpiration cooling system . 13
Figure 5-1: Surface energy balance. 17
Figure 5-2: Galileo entry probe. . 20
Figure 5-3: Physical model and phenomena considered in material response analysis . 20
Figure 5-4: Temperature history at interfaces. . 22
Figure 5-5: Comparison of mass loss fluxes. . 22
Figure 6-1: Worst case peak predicted surface temperatures. [K] for STS-1. From Dotts
et al. (1983) [5]. . 25
Figure 6-2: Worst case peak predicted structure temperatures. [K] for STS-1. From
Dotts et al. (1983) [5]. . 25
Figure 6-3: Thermal protection subsystems. From Dotts et al. (1983) [5] . 26
Figure 6-4: RCC system components. From Curry et al. (1983) [3]. . 27
Figure 6-5: Nose cap system components. From Curry et al. (1983) [3]. . 27
Figure 6-6: Wing leading-edge system components. From Curry et al. (1983) [3]. . 28
Figure 6-7: Tile attachment and gap filler configuration. From Dotts et al. (1983) [5]. . 29
Figure 6-8: Nose cap RCC surface comparison between prediction and flight data.
From Curry et al. (1983) [3] . 30
Figure 6-9: Nose cap access door tile surface comparison between prediction and flight
data. From Curry et al. (1983) [3]. . 30
Figure 6-10: Wing leading-edge panel (stagnation area). Comparison between
prediction and flight data. From Curry et al. (1983) [3]. . 31
Figure 6-11: STS-1 flight data analysis comparison for lower mid-fuselage location.
From Dotts et al. (1983) [3]. 31
Figure 6-12: STS-1 flight data analysis comparison for lower wing location. From Dotts
et al. (1983) [3] . 32
Figure 6-13: STS-1 flight data analysis comparison for lower inboard elevon location.
From Dotts et al. (1983) [3]. 32
Figure 6-14: STS-1 flight data analysis comparison for lower mid-fuselage side
location. From Dotts et al. (1983) [3]. . 33
Figure 6-15: Comparison of STS-2 data with analytical predictions. From Normal et al.
(1983) [11]. . 33
Figure 6-16: Comparison of STS-2 data with analytical predictions. From Normal et al.
(1983) [11]. . 34
Figure 6-17: Comparison of STS-2 data with analytical predictions. From Normal et al.
(1983) [11]. . 34
Figure 6-18: In-depth comparison of STS-2 data with analytical predictions for
maximum temperatures. From Normal et al. (1983) [11]. . 35
Figure 6-19: X-39 TPS Configuration . 36
Figure 6-20: X-38 Reference Heating . 36
Figure 6-21: CMC Side Panels together with lower CMC Chin Panel . 37
Figure 6-22: Stand-off Position and Global Design . 38
Figure 6-23: Stand-off Positions and Global Design . 39
Figure 6-24: Max. Pressure Load . 40
Figure 6-25: Max. Thermal Load at Panel Surface . 40
Figure 6-26: Nose Skirt Assembly with Insulation Blankets . 41
Figure 6-27: Max. and min. Heat flux time lines applied on the NSK. 41
Figure 6-28: Simplified description of heat transfer modes within the nose skirt
assembly. . 42
Figure 6-29: Temperature distribution over a NSK side panel at t = 1100s. . 44
Figure 6-30: Carrier Panel TPS Design . 45
Figure 6-31: X-38 Aeroshell Panel and Blanket Distribution . 46
Figure 6-32: X-38 Parafoil System . 46
Figure 6-33: Parafoil Line Routing and Acreage Blankets . 46
Figure 6-34: FEI-450 Blanket equipped with Gray FEI-1000High Emittance Coating . 47
Figure 6-35: Typical look of FEI-650 and Blanket with Gray High Emittance . 47
Figure 6-36: Allocation of Blanket Types to the X-38 Lee-Side Surface . 49
Figure 6-37: Qualification Test Sequence for X-38 . 50
Figure 6-38: Parameters and Results of the Qualification Tests . 50
Figure 6-39: Computer controlled sewing of FEI blankets . 52
Figure 6-40: FEI-1000 blankets of the Forward Fuselage . 52
Figure 6-41: FEI Blankets Integrated on the X-38 V-201 . 53

European Foreword
This document (CEN/CLC/TR 17603-31-16:2021) has been prepared by Technical Committee
CEN/CLC/JTC 5 “Space”, the secretariat of which is held by DIN.
It is highlighted that this technical report does not contain any requirement but only collection of data
or descriptions and guidelines about how to organize and perform the work in support of EN 16603-
31.
This Technical report (TR 17603-31-16:2021) originates from ECSS-E-HB-31-01 Part 16A.
Attention is drawn to the possibility that some of the elements of this document may be the subject of
patent rights. CEN [and/or CENELEC] shall not be held responsible for identifying any or all such
patent rights.
This document has been prepared under a mandate given to CEN by the European Commission and
the European Free Trade Association.
This document has been developed to cover specifically space systems and has therefore precedence
over any TR covering the same scope but with a wider domain of applicability (e.g.: aerospace).
Scope
The thermal protection system (TPS) of a space vehicle ensures the structural integrity of the surface of
the craft and maintains the correct internal temperatures (for crew, electronic equipment, etc.) when
the vehicle is under the severe thermal loads of re-entry. These loads are characterised by very large
heat fluxes over the relatively short period of re-entry.
The design of thermal protection systems for re-entry vehicles is very complex due to the number and
complexity of phenomena involved: the flow around the vehicle is hypersonic, tridimensional and
reactive, and its interaction with the vehicle’s surface may induce chemical reactions which are not
fully understood.
Two TPS concepts for re-entry vehicles, ablative and radiative are examined and there is also an
anlyisis of existing systems using them.

The Thermal design handbook is published in 16 Parts
TR 17603-31-01 Thermal design handbook – Part 1: View factors
TR 17603-31-02 Thermal design handbook – Part 2: Holes, Grooves and Cavities
TR 17603-31-03 Thermal design handbook – Part 3: Spacecraft Surface Temperature
TR 17603-31-04 Thermal design handbook – Part 4: Conductive Heat Transfer
TR 17603-31-05 Thermal design handbook – Part 5: Structural Materials: Metallic and
Composite
TR 17603-31-06 Thermal design handbook – Part 6: Thermal Control Surfaces
TR 17603-31-07 Thermal design handbook – Part 7: Insulations
TR 17603-31-08 Thermal design handbook – Part 8: Heat Pipes
TR 17603-31-09 Thermal design handbook – Part 9: Radiators
TR 17603-31-10 Thermal design handbook – Part 10: Phase – Change Capacitors
TR 17603-31-11 Thermal design handbook – Part 11: Electrical Heating
TR 17603-31-12 Thermal design handbook – Part 12: Louvers
TR 17603-31-13 Thermal design handbook – Part 13: Fluid Loops
TR 17603-31-14 Thermal design handbook – Part 14: Cryogenic Cooling
TR 17603-31-15 Thermal design handbook – Part 15: Existing Satellites
TR 17603-31-16 Thermal design handbook – Part 16: Thermal Protection System

References
EN Reference Reference in text Title
EN 16603-00-01 ECSS-S-ST-00-01 ECSS System - Glossary of terms

All other references made to publications in this Part are listed, alphabetically, in the Bibliography.
Terms, definitions and symbols
3.1 Terms and definitions
For the purpose of this Standard, the terms and definitions given in ECSS-S-ST-00-01 apply.
3.2 Abbreviated terms
The following abbreviated terms are defined and used within this Standard.
computer aided design
CAD
computational fluid dynamics
CFD
ceramics matrix composite
CMC
carbon reinforced silicon carbide
C/SiC
flexible external insulation
FEI
flexible reusable surface insulation
FRSI
high temperature insulation
HTI
high-temperature reusable surface insulation
HRSI
internal flexible insulation
IFI
low-temperature reusable surface insulation
LRSI
reinforced carbon-carbon
RCC
reusable surface insulation
RSI
strain isolation pad
SIP
structural outer mold line
SOML
TPS outer mold line
TOML
thermal protection system
TPS
Introduction
4.1 General
The thermal protection system (TPS) of a space vehicle consists of those elements needed to protect
the structural integrity of the vehicle’s surface and maintain the appropriate internal temperatures (for
crew, electronic equipment, etc.) when the vehicle is under the severe thermal loads of re-entry. These
loads are mainly characterised by very large heat fluxes during relatively short times.
The heat fluxes acting on the TPS are so large because of the great speeds of re-entry vehicles. The
velocity-altitude map for the Space Shuttle is represented in Figure 4-1.

Figure 4-1: Velocity-altitude map for the Space Shuttle. Lifting re-entry from orbit.
The heat fluxes and the time of re-entry are basically determined by the re-entry orbit. These orbits are
designed so that the vehicle is captured by the planet and the payload is not damaged by the
accelerations; these factors greatly restrict the number of valid trajectories. However, for lifting
vehicles which can be manoeuvred those restrictions are alleviated, and re-entry trajectories, other
than ballistic, can be achieved. In Figure 4-2 the heat fluxes and re-entry times for different trajectories
are summarised.
Figure 4-2: Summary of re-entry trajectories. From East (1991) [6].
The design of thermal protection systems for re-entry vehicles is a very complex problem due to the
number and complexity of phenomena involved. It suffices to mention here that the flow around the
vehicle is hypersonic, tridimensional and reactive, and its interaction with the vehicle’s surface may
induce chemical reactions which are not fully understood.
4.2 Classification of thermal protection systems
Generally speaking the TPS consists of a material system (shield and/or load carrying member)
operating on a given heat dissipation principle. There are several TPS concepts for re-entry vehicles
(Hurwicz & Rogan (1973a) [9]):
• Ablative thermal protection
• Radiative thermal protection
• Heat sinks
• Transpiration cooling
ABLATIVE SYSTEMS
Ablative systems operate dissipating the incident thermal energy through the loss of material: these
systems lose mass as a consequence of the ablation of the external surface material. They have good
thermal characteristics since phase changes absorb a large amount of energy. These systems are not
reusable. See Figure 4-3 for a sketch of an ablative system.
Figure 4-3: Sketch of an ablative thermal protection system.
The ablation process is quite complex and is described in some detail in clause 5.2. One important
consequence of the analysis of these systems is that their efficiency is particularlysensitive to material
performance. Therefore, it is necessary to treat the subject of materials in detail. In the absence of a
universally acceptable ablative material a wide variety of ablative compositions and constructions
have been produced, usually tailored to satisfy the requirements of a specific vehicle for a specific
mission. A detailed description of ablative materials is given in clause 5.3.

RADIATIVE SYSTEMS
Radiative systems operate re-emitting by radiation the energy received from the surrounding
environment. They are composed of two layers: an outer layer which consists of a material that can
stand the radiation equilibrium temperature and an inner layer which insulates the outer layer from
the structure in order to minimise the heat flow between the two, see Figure 4-4.

Figure 4-4: Sketch of a radiative thermal protection system.
It will be seen in clause 6.1 that the effectiveness of a radiative system increases very rapidly with
increasing surface temperatureand surface emissivity. Consequently, the primary development efforts
have been concerned with the improvement of high emissivity, high temperature coatings, and with
increasing the material service temperatures (including that of the internal insulation). A detailed
description of materials used in radiative systems is given in clause 6.2.
These systems can be designed including a cooling subsystem: this is a fluid loop where the working
fluid transports heat from the areas where the heat flux is stronger to those where the heat flux is
weaker. The actual mechanism for heat transport can be the same as in heat pipes, the fluid is
vaporised in areas of higher temperatures, and it is condensed in areas of lower temperatures.
However, even though the characteristics of these systems are good, they are not used in practice.

HEAT SINK THERMAL PROTECTION SYSTEM
A heat sink is the simplest type of absorptive thermal protection system. It was used in the design of
the early re-entry vehicles (e.g. the first two manned Mercury vehicles).
These systems are composed of an outer layer, comparatively thick, which consists of a material of
high conductivity and capacitance. The function of this layer is to absorb the heat input. Since the
material heats up, the storage capability is limited by the melting temperature.
Its use is limited to relatively low heating rates and therefore may not be practical for the high heat
loads encountered in short re-entry times.
Heat sinks have the advantages of simplicity, dependability, and for reusable vehicles, ease of
refurbishment. Their outstanding disadvantage is their low efficiency, this would cause a heat sink
sized to satisfy most current re-entry missions to be excessively heavy.
Materials commonly used as heat sinks are
• beryllium
• beryllium oxide (beryllia)
• copper.
Graphite has many desirable heat sink characteristics, but begins to oxidise at temperatures far below
those required for best efficiency.

TRANSPIRATION COOLING
Transpiration systems are systems where fluid is injected through a porous medium into the
boundary layer. The structure is maintained cool by two basic mechanisms: heat is conducted to the
coolant as it flows through the structure, and as the coolant is ejected out the surface it reduces the
surface heat transfer rate by cooling and thickening the boundary layer. See Figure 4-5 for a sketch.

Figure 4-5: Sketch of a transpiration thermal protection system.
In some applications, the shape change caused by the surface recession of an ablating surface is not
acceptable for aerodynamic performance reasons. In such cases, if the environment is too severe for
radiative or heat sink systems, transpiration cooling may be the only practical solution. This TPS
makes possible performance in environments that could not otherwise be withstood. However, its
mechanical complexity (see Figure 4-6), with the associated reliability problems, tend to limit its use.
Figure 4-6: Typical transpiration cooling system
For re-entry application, the most acceptable coolants are:
• H2O
• NH
• CF4
• CO
Ablative systems
5.1 General
The ablation mechanism for thermal protection is based on the sublimation, melting or pyrolysis of
the heat shield and the removal of the products by the outer stream. The great amount of energy
absorbed in phase transition reduces the heat fluxes to the structure of the vehicle. This method has
been widely used in most of non–reusable entry vehicles, for its simplicity and its high performance. It
has been used in planetary probes, ballistic missiles and space capsules.
The methods of analysing the various heat shield materials vary depending on the melt temperature
and oxidation chemistry. These materials may be classified as
a. Oxidation controlled. The melting temperature is greater than the radiative equilibrium
temperature calculated for the convective heat transfer rate.
b. Simple sublimers. When melting temperature is lower than the radiative equilibrium
temperature.
c. Pyrolytic ablators. The material decomposition into pyrolysis gas and char occurs in depth.
Despite the above classification almost all heat shields are made with carbon based materials. This fact
is due to special characteristic combination of very desirable properties as good heat sink, high melt
temperature, large heat of sublimation, good radiation properties, and from the structural point of
view low dilatation coefficient.
5.2 Ablative materials
Materials commonly used can be classified as follows:
Composites:
Carbon phenolic:
High strength charring ablator with high ablation temperature.
Used in high performance re-entry vehicles.
It is the current heat shield material choice.
Used in the Galileo probe (clause 5.4.1).
Silica phenolic:
High strength charring ablator.
Used in high performance re-entry vehicles.
Selected for the Huygens probe.
Phenolic nylon:
High heat shielding capability.
It has the disadvantage of high erosion rate.
Limited capability to accommodate high heat loads.
Used in the Galileo probe aft shield (clause 5.4.1).

Ceramics:
Graphite:
Used in high heating rate areas.
It has disadvantages as brittleness and low resistance to thermal stress.
ATJ graphite has high strength properties.

Metals:
Tungsten:
Refractory metal.
Used as a porous matrix infiltrated with copper or silver.
It has good mechanical properties but low thermal performance.

Elastomers:
Silicone polymers:
Used reinforced by ceramic microspheres.
Used in low shear and low pressure regions, and in low to moderate heat flux zones.
It was used in the Gemini project.

Plastics:
Teflon:
Low temperature ablator with moderate efficiency and high ablation rates.
It has been used in ballistic missiles.

AVCOAT 5026.
Ablator used in the Apollo capsules.
5.3 Basic analysis
The methods for predicting surface degradation or recession of ablative thermal protection systems
have been purely empirical or semitheoretical. The semitheoretical methods are based on simplifying
assumptions that will be explained later. These simplifications are due to the extremely complicated
physico-chemical phenomena involved in the ablation process. It includes phase change, non–
equilibrium effects, multiphase flow, high thermal radiation environment, three dimensional
hypersonic flow. The design of high mass heat shields has made it necessary accurate theoretical
models for ablation. An accurate design of these heat shields is critical due to the high mass of ablator
necessary for protecting the vehicles. It can be noticed that heavy heat shields may make missions not
feasible. An example of accurate theoretical design is the Galileo Jovian probe, it has been designed
using numerical detailed flow field predictions, despite the uncertainties on the Jupiter atmosphere
composition.
5.3.1 Surface equilibrium
To establish the energy flux to the re-entry vehicle, the energy transfer mechanisms between the
boundary layer and the ablating surface of the vehicle must be stated. Despite the fact that the ablation
problem is non–steady, the assumption of steady state will be made. This approximation is rather
accurate for engineering purposes. It is assumed, also, that the fluid flow near the surface is governed
by the equations of the boundary layer with mass injection. In this analysis (Hurwicz & Rogan (1973b)
[10]) the conservation of mass and energy are applied to a thin control volume (since this is a moving
control volume the Reynolds transport theorem should be taken into account).
First the mass balance is considered. The mass flux leaving the control volume through the gas side is
 
due to gaseous species, that are denoted by m , plus solid and liquid species, denoted by m .The
w r

term m is due to sublimation of the heat shield and products from pyrolysis reactions that occur in
w
the heat shield. The term m is due to mechanical erosion of liquid and solid products, the solid
r
erosion is due to spallation of solid particles and to the surface stress due to percolation and
sublimation, the liquid products are due to heat shield melting. This term, m , should be reduced
r
because this mass does not sublime, thus reducing the performance of the heat shield. The mass flux
entering the control volume through the solid side is due to the products of pyrolysis reactions that
occur in the heat shield, and to the solid mass that enters due to the movement of the control volume.
 
The first flux is denoted by m and the second one by m . The continuity equation yields
g c
S S
   
m + m =m +m
[5-1]
w r c g
S S
with
 ( )  ( )
m = m , m = m
∑ ′ ∑ ′′
c k w g k w
S S [5-2]
k′ k′′
where k’ denotes summation over all the solid species entering the control volume, and k" denotes
summation over the gaseous species.
The energy balance yields
q = q − m h − m h+ m h +
s w w w r r c c
S S
[5-3]
 
+m h −εσT +α q
g g w w w rad
S S
where the meaning of each term is explained below.
q is the energy flux to the solid, it is positive when the energy flux is to the solid.
s
q is the energy flux due to the aerodynamic heat transfer to the ablator from the boundary layer.
w
 is the energy flux associated to the gaseous mass coming into the boundary layer from the solid
m h
w w
boundary.
h = Y(h)
w ∑ i i
w [5-4]
i
with
T
w
o
(h) = c dT+ h
[5-5]
i p i
w
∫ i
T
o
o
where Yi is the mass fraction of the i-th gaseous species and hi is the formation enthalpy at a reference
state (po,To).

m h is the energy flux associated to the solid and liquid removal from the solid boundary.
r r
m h is the energy flux corresponding to the solid mass flux associated to the recession of the heat
c c
S S
shield.
m h = (m )(h )
c c ∑ k′ k′
w w
S S [5-6]
′
k

m h is the energy flux corresponding to the gas mass flux associated to pyrolisis reactions.
g g
S S
 ( )( )
m h = m h
∑ ′′ ′′
g g k w k w
S S [5-7]
k′′
εwσTw is the energy flux radiated by the solid surface, where εw is the surface emissivity and σ is the
Stefan-Boltzman constant.
 
α q is the energy flux absorbed from the fluid, where αw is the absorptance; q should be
w rad rad
obtained from the shock layer solution.
In Figure 5-1 the surface energy balance is sketched, where the different quantities appearing in the
equation are positive if the fluxes have the sense shown in the figure.

Figure 5-1: Surface energy balance
In order to make use of the previous analysis, the different terms in the energy balance need to be
evaluated. This is a difficult interdisciplinary task. As an example some results for the aerodynamic
heating from the boundary layer,  , are summarised here.
q
w
For air, Fay and Riddle (quoted in Anderson (1989) [1]) have correlated many experimental results
obtaining different correlations for the heat flux at the stagnation point of a spherical nose. For
equilibrium boundary layer the heat flux is

0,4 0,1
−0,6

q = 0,76Pr (ρµ ) (ρ µ )
w e e w w
[5-8]
 
du h
0,52
e c
(l − h ) 1+(Le −1)
e w 
ds l
 e
For a frozen boundary layer with a fully catalytic wall

−0,6 0,4 0,1

q = 0,76Pr (ρµ ) (ρ µ )
w e e w w
[5-9]
 
du h
0,63
e c
(l − h ) 1+(Le −1)
e w 
ds l
 e
For a frozen boundary layer with a noncatalytic wall

 
du h
0,4 0,1
−0,6
e c

q = 0,76Pr (ρµ ) (ρ µ ) 1−
[5-10]
w e e w w  
ds l
 e
where
µ c
e p
e
Pr= Prandtl> number
[5-11]
k
e
ρ c D
e p
e
Le= Lewis> number
[5-12]
k
e
ρe density outside the boundary layer
µe viscosity outside the boundary layer
ke thermal conductivity outside the boundary layer
cpe specific heat at constant pressure outside the boundary layer
D diffusion coefficient
ρw density at the wall
µw viscosity at the wall
Ie total enthalpy of the stream outside the boundary layer
T
e 2
U
e
I = c dT+
[5-13]
e p
∫
T
o
where Ue is the velocity outside the boundary layer
hw enthalpy at the wall
T
e
h = c dT
[5-14]
w p
∫
T
o
o
( )
h = Y h
∑
c i e i
[5-15]
i
o
where (Yi)e is the mass fraction of the i-th species outside the boundary layer and hi is the formation
enthalpy of the i-th species at the reference state
and the stagnation point velocity gradient is given by Newtonian theory (Anderson (1989) [1]) as

du 1 2(p − p )
e e ∞
=
[5-16]
ds R ρ
e
where pe is the pressure outside the boundary layer, p∞ the pressure far ahead of the spherical nose,
and R is the local radius of curvature of the body at the stagnation point.
5.4 Existing systems
5.4.1 Galileo probe
The mission of the Galileo entry probe is to descend through Jupiter’s atmosphere, and it is
instrumented so that the planet’s atmospheric structure, composition, and radiation balance can be
measured. The probe is illustrated in Figure 5-2. The main components are the descent module, which
contains the scientific measurement package, the aeroshell structure and the thermal protection
system. The TPS consists of two heat shields, one for the forebody and another for the afterbody. The
forebody shield consists of two parts, the spherical nose composed of chopped-molded carbon
phenolic, and the frustrum composed of 30° tape-wrapped carbon phenolic. The probe aft shield is
made of phenolic nylon. The ablative material is bonded to the aeroshell which is made of aluminium.
Figure 5-2: Galileo entry probe.
In this clause the study on the ablative thermal protection system made by Green and Davy (1982) [8]
will be described. This study corresponds to a 310 kg probe (other studies for a 242 kg and a 290 kg
probes are also available). The actual probe mass is 339 kg.
5.4.1.1 Heat shield analysis
The analysis of the heat-shield is divided into four parts which are described next (see Figure 5-3).

Figure 5-3: Physical model and phenomena considered in material response
analysis
1. Specifying the entry heating environment.
2. Modelling the in-depth phenomena:
o conduction
o receding surface
o decomposition
o pyrolysis
o storage
3. Modelling the surface phenomena:
o ablative heating
o reradiation
o equilibrium chemistry
o ablation
o elemental mass diffusion
o char removal
o pyrolysis
4. Specifying the interface constraints necessary to size the shield. In this case they are the
following:
o bond temperature less than 644 K
o aeroshell temperature less than 589 K
o insulated back surface of the aeroshell
This analysis also requires knowledge of material properties for the carbon-phenolic (including the
char), the bond and the aeroshell.
5.4.1.2 Results
The different models (including material-property models) used in the analysis are described in Green
and Davy (1982) [8]. In this clause some results will be presented. In Figure 5-4 the response at the
three key interfaces is shown; results for two models, quasi-steady and transient, are presented. The
difference in maximum temperature between both models is about 5%. The mass loss fluxes are
presented in Figure 5-5. The time-integrated mass loss indicates that the quasi-steady result is about
20% higher. It can be noted that these differences are actually due to the use of different material-
property models.
Figure 5-4: Temperature history at interfaces.

Figure 5-5: Comparison of mass loss fluxes.
With respect to the total mass of ablator, different material-property models give different results,
ranging from 117 kg (of which 93 kg is lost) to 139 kg (of which 120 kg is lost). The principal difference
in these models is the value of the surface reflectance. The heat-shield mass requirement is so large
(around 40% of the total mass) that a heat-shield design based on a conservative safety margin,
although desirable, is not possible, except at the expense of other subsystems, because the total probe
mass is constrained by launch requirement limits. Hence, an accurate heat-shield design is crucial to
the success of the mission.
Radiative systems
6.1 General
In this second class of thermal protection systems there is no mass loss; these systems can be reused.
The thermal protection is achieved by re-emitting in form of thermal radiation the energy received by
the solid walls.
In these systems energy absorption is small, and thus the surface energy balance can be expressed as

q =σε T
[6-1]
w w w
where Tw is the radiation equilibrium temperature and is the maximum temperature the surface can
experience for a given heat flux. The radiative system is therefore limited to a maximum heat rate (by
virtue of its operating temperature limit) rather than a total heat input as an ablator may be. Once an
operating temperature limit is defined, the maximum heat flux is correspondingly defined and the
system can operate indefinitely at this condition. The penalty associated with increased operating time
is the amount of insulation required between the heat shield and inner surface (either the load
carrying structure or the payload).
As evidenced by the equation above, the effectiveness of a radiative system increases very rapidly
with increasing surface temperature. The surface emissivity, ε, also exerts an important, although
lesser, effect on the amount of energy which can be accommodated.
6.2 Radiative materials
External insulators commonly used can be divided in two classes: rigid and flexible.

RIGID INSULATORS
Rigid insulators can adopt different shapes: tiles, shingles, shells and boxes. Materials used to make
them can be classified as follows:
Composites:
• Carbon carbon
• Reinforced carbon carbon (see clause 6.3.1.1)
• Carbon/Silicon carbide
Ceramics:
• Sinterized alumina/silica fibres
• Sinterized high-purity silica fibres
This type of insulation is used to protect areas exposed to the highest temperatures.

FLEXIBLE INSULATORS
Flexible insulators are blankets of different materials:
• Silica fibre
• Glass fibre
• Alumina/silica fibre
• Alumina/borosilicate fibre
• Nomex fibre
• Alumina fibre plated with rhodium
• Nylon
These materials are processed into fleeces, felts or threads, which then form the blankets.
6.3 Existing systems
6.3.1 Space shuttle
In this clause the thermal protection system of the Space Shuttle Orbiter will be described. This system
is designed so that it can be used 100 times or more without major refurbishment. The surface of the
Shuttle is subjected to temperatures ranging from 600 K to 1750 K, as shown in Figure 6-1, and the TPS
is designed to keep the vehicle structure at the temperatures shown in Figure 6-2.
Figure 6-1: Worst case peak predicted surface temperatures. [K] for STS-1. From
Dotts et al. (1983) [5].
Figure 6-2: Worst case peak predicted structure temperatures. [K] for STS-1. From
Dotts et al. (1983) [5].
TPS multimission capability is obtained by a combination of four different material configurations,
which are optimised for specific operational temperature ranges.
1. Temperatures greater than 1500 K. The thermal protection subsystem for this region
consists of coated reinforced carbon-carbon (RCC).
2. Temperatures from 900 K to 1500 K. This region is protected with a high-temperature
reusable surface insulation (HRSI).
3. Temperatures from 650 K to 900 K. In this region the protection is a low-temperature
reusable surface insulation (LRSI).
4. Temperatures lower than 650 K. This region is protected with a flexible reusable surface
insulation (FRSI).
These four thermal protection subsystems, see Figure 6-3, will be described next.

Figure 6-3: Thermal protection subsystems. From Dotts et al. (1983) [5]
6.3.1.1 Description of thermal protection subsystems
REINFORCED CARBON-CARBON
This subsystem consists of the RCC nose cap and RCC wing leading-edge panels, the metallic
attachments to the Orbiter structure, the internal insulation system, thermal barriers, and the interface
tiles between the RCC and reusable surface insulation (RSI).
The RCC panels function as the airfoil shape at temperatures exceeding 1500 K. The wing leading
edge consists of 44 RCC panels/T–seals (22 on each wing), whereas the nose cap is a RCC monoconic
shell with associated seals. Figure 6-4, Figure 6-5 and Figure 6-6 show the major components of the
nose cap and wing leading-edge systems.
Figure 6-4: RCC system components. From Curry et al. (1983) [3].

Figure 6-5: Nose cap system components. From Curry et al. (1983) [3].
Figure 6-6: Wing leading-edge system components. From Curry et al. (1983) [3].
The RCC material is a hard carbon structure highly resistant to fatigue loads, possessing reasonable
strength and a low coefficient of thermal expansion which provides it with excellent resistance to
thermal stresses and shock. Since the RCC parts form a hollow shell, they promote internal cross
radiation from the hot stagnation region to cooler areas, thus reducing stagnation temperatures and
thermal gradients around the shell. The operational temperature of the RCC gets up to about 1900 K.
Since RCC is not an insulator, the adjacent aluminium and the
...