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

Status
Published
Public Enquiry End Date
26-May-2021
Publication Date
23-Aug-2021
Technical Committee
Current Stage
6060 - National Implementation/Publication (Adopted Project)
Start Date
19-Aug-2021
Due Date
24-Oct-2021
Completion Date
24-Aug-2021

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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-
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).

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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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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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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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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
• NH
• CF4
• CO
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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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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

notice and shall not be referred to as a Technical Report.
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. FprCEN/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 (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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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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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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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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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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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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