SIST-TP CEN/CLC/TR 17603-31-16:2021
(Main)Space Engineering - Thermal design handbook - Part 16: Thermal Protection System
Space Engineering - Thermal design handbook - Part 16: Thermal Protection System
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 Part 1A Thermal design handbook – Part 1: View factors
TR 17603-31-01 Part 2A Thermal design handbook – Part 2: Holes, Grooves and Cavities
TR 17603-31-01 Part 3A Thermal design handbook – Part 3: Spacecraft Surface Temperature
TR 17603-31-01 Part 4A Thermal design handbook – Part 4: Conductive Heat Transfer
TR 17603-31-01 Part 5A Thermal design handbook – Part 5: Structural Materials: Metallic and Composite
TR 17603-31-01 Part 6A Thermal design handbook – Part 6: Thermal Control Surfaces
TR 17603-31-01 Part 7A Thermal design handbook – Part 7: Insulations
TR 17603-31-01 Part 8A Thermal design handbook – Part 8: Heat Pipes
TR 17603-31-01 Part 9A Thermal design handbook – Part 9: Radiators
TR 17603-31-01 Part 10A Thermal design handbook – Part 10: Phase – Change Capacitors
TR 17603-31-01 Part 11A Thermal design handbook – Part 11: Electrical Heating
TR 17603-31-01 Part 12A Thermal design handbook – Part 12: Louvers
TR 17603-31-01 Part 13A Thermal design handbook – Part 13: Fluid Loops
TR 17603-31-01 Part 14A Thermal design handbook – Part 14: Cryogenic Cooling
TR 17603-31-01 Part 15A Thermal design handbook – Part 15: Existing Satellites
TR 17603-31-01 Part 16A 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
Vesoljska tehnika - Priročnik o toplotni zasnovi - 16. del: Sistem toplotne zaščite
General Information
Buy Standard
Standards Content (Sample)
SLOVENSKI STANDARD
SIST-TP CEN/CLC/TR 17603-31-16:2021
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
SIST-TP CEN/CLC/TR 17603-31-16:2021 en,fr,de
2003-01.Slovenski inštitut za standardizacijo. Razmnoževanje celote ali delov tega standarda ni dovoljeno.
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TECHNICAL REPORT
CEN/CLC/TR 17603-31-
16
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.
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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
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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
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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
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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).
5
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1
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
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2
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.
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3
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
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4
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.
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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.
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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.
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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.
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Figure 4-6: Typical transpiration cooling system
For re-entry application, the most acceptable coolants are:
• H2O
3
• NH
• CF4
2
• CO
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5
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.
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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
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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 spec
...
SLOVENSKI STANDARD
kSIST-TP FprCEN/CLC/TR 17603-31-16:2021
01-maj-2021
Vesoljska tehnika - Priročnik za toplotno zasnovo - 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: FprCEN/CLC/TR 17603-31-16
ICS:
49.140 Vesoljski sistemi in operacije Space systems and
operations
kSIST-TP FprCEN/CLC/TR 17603-31- en,fr,de
16:2021
2003-01.Slovenski inštitut za standardizacijo. Razmnoževanje celote ali delov tega standarda ni dovoljeno.
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kSIST-TP FprCEN/CLC/TR 17603-31-16:2021
TECHNICAL REPORT
FINAL DRAFT
FprCEN/CLC/TR 17603-
RAPPORT TECHNIQUE
31-16
TECHNISCHER BERICHT
March 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: Wärmeschutzsystem
spatiaux
This draft Technical Report is submitted to CEN members for Vote. 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.
Recipients of this draft are invited to submit, with their comments, notification of any relevant patent rights of which they are
aware and to provide supporting documentation.
Warning : This document is not a Technical Report. It is distributed for review and comments. It is subject to change without
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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
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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
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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
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European Foreword
This document (FprCEN/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.
This document is currently submitted to the Vote on TR.
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 (FprCEN/CLC/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).
This document is currently submitted to the CEN CONSULTATION.
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1
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
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2
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.
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3
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
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4
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.
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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.
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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.
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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.
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Figure 4-6: Typical transpiration cooling system
For re-entry application, the most acceptable coolants are:
H2O
NH3
CF4
CO2
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5
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.
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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
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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 assumpt
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