SIST-TP CEN/CLC/TR 17603-31-15:2021

Space Engineering - Thermal design handbook - Part 15: Existing Satellites

In this Part 15, existing satellites are described and examined from a thermal control and design view. The thermal control requirements are given and an assessment is made of the thermal control systems used against performance for each satellite.
The Thermal design handbook is published in 16 Parts
TR 17603-31-01 Part 1    Thermal design handbook – Part 1: View factors
TR 17603-31-01 Part 2    Thermal design handbook – Part 2: Holes, Grooves and Cavities
TR 17603-31-01 Part 3    Thermal design handbook – Part 3: Spacecraft Surface Temperature
TR 17603-31-01 Part 4    Thermal design handbook – Part 4: Conductive Heat Transfer
TR 17603-31-01 Part 5    Thermal design handbook – Part 5: Structural Materials: Metallic and Composite
TR 17603-31-01 Part 6    Thermal design handbook – Part 6: Thermal Control Surfaces
TR 17603-31-01 Part 7    Thermal design handbook – Part 7: Insulations
TR 17603-31-01 Part 8    Thermal design handbook – Part 8: Heat Pipes
TR 17603-31-01 Part 9    Thermal design handbook – Part 9: Radiators
TR 17603-31-01 Part 10    Thermal design handbook – Part 10: Phase – Change Capacitors
TR 17603-31-01 Part 11    Thermal design handbook – Part 11: Electrical Heating
TR 17603-31-01 Part 12    Thermal design handbook – Part 12: Louvers
TR 17603-31-01 Part 13    Thermal design handbook – Part 13: Fluid Loops
TR 17603-31-01 Part 14    Thermal design handbook – Part 14: Cryogenic Cooling
TR 17603-31-01 Part 15    Thermal design handbook – Part 15: Existing Satellites
TR 17603-31-01 Part 16    Thermal design handbook – Part 16: Thermal Protection System

Raumfahrttechnik - Handbuch für thermisches Design - Teil 15: Bestehende Satelliten

Ingénierie spatiale - Manuel de conception thermique - Partie 15: Véhicules spatiaux existants

Vesoljska tehnika - Priročnik o toplotni zasnovi - 15. del: Obstoječi sateliti

General Information

• Status: Published
• Public Enquiry End Date: 26-May-2021
• Publication Date: 23-Aug-2021
• ICS: 49.140 - Space systems and operations
• Technical Committee: I13 - Imaginarni 13
• Current Stage: 6060 - National Implementation/Publication (Adopted Project)
• Start Date: 19-Aug-2021
• Due Date: 24-Oct-2021
• Completion Date: 24-Aug-2021
• Mandate: M/496 - Mandate addressed to CEN, CENELEC and ETSI to develop standardisation regarding space industry (Phase 3 of the process)
• Ref Project: CEN/CLC/TR 17603-31-15:2021 - Space Engineering - Thermal design handbook - Part 15: Existing Satellites

Overview

SIST-TP CEN/CLC/TR 17603-31-15:2021 - Space Engineering: Thermal design handbook - Part 15: Existing Satellites documents the thermal design and thermal control experience of flown spacecraft. This technical report reviews a selection of existing satellites, describing missions, main subsystems, orbit characteristics, thermal design requirements, and the thermal control systems used. For each spacecraft the report presents design trade‑offs, thermal control of key components, verification and thermal test information, and estimated and measured on‑orbit performance.

Key topics and technical requirements

• Scope and structure: Definitions, symbols, and references; case studies of multiple historical satellites (e.g., IUE, OTS, Landsat D, IRAS, SPOT, Olympus‑1, ERS‑1) with consistent thermal reporting.
• Thermal design requirements: Functional temperature limits, thermal interfaces and orbital constraints that drive design margins and control strategies.
• Design trade‑offs and architectures: Passive vs active thermal control choices, coatings/insulations, heaters, fluid loops, and radiators as applied on existing platforms.
• Component‑level control: Thermal management approaches for batteries, electronics, payloads, cryogenic systems and propulsion tanks.
• Verification and test data: Thermal vacuum/thermal balance tests, ground vs on‑orbit performance comparison, histograms and trend analyses to validate thermal models.
• Performance assessment: Measured on‑orbit temperatures and long‑term trends used to evaluate system effectiveness and lessons learned.

Practical applications and target users

This Part 15 is valuable for:

• Spacecraft thermal engineers performing design reviews, model validation and heritage assessments.
• Systems and mission engineers defining thermal requirements and interfaces during concept and trade‑study phases.
• Test engineers planning thermal vacuum and thermal balance tests using historical test configurations and results as references.
• Program managers and procurements seeking evidence of proven thermal approaches and risk‑reduction strategies.
• Academic and research teams studying historical thermal control performance and spacecraft thermal reliability.

Use cases include heritage benchmarking, failure analysis, retrofit planning for existing satellites, and educating teams on practical thermal control solutions validated in flight.

Related standards (Thermal design handbook series)

Part 15 is one of 16 parts in the TR 17603‑31 series covering foundational thermal topics and technologies:

• Parts 1–4 (View factors, Holes/Grooves/Cavities, Surface Temperature, Conductive Heat Transfer)
• Parts 5–14 (Materials, Surfaces, Insulations, Heat Pipes, Radiators, Phase‑Change, Electrical Heating, Louvers, Fluid Loops, Cryogenic Cooling)
• Part 16 (Thermal Protection System)

This Part 15 complements those parts by providing real‑world spacecraft thermal design examples and measured performance-making it a practical resource for anyone involved in spacecraft thermal engineering and verification.

Standards Content (Sample)

SLOVENSKI STANDARD
01-oktober-2021
Vesoljska tehnika - Priročnik o toplotni zasnovi - 15. del: Obstoječi sateliti
Space Engineering - Thermal design handbook - Part 15: Existing Satellites
Raumfahrttechnik - Handbuch für thermisches Design - Teil 15: Bestehende Satelliten
Ingénierie spatiale - Manuel de conception thermique - Partie 15: Véhicules spatiaux
existants
Ta slovenski standard je istoveten z: CEN/CLC/TR 17063-31-15:2021
ICS:
49.140 Vesoljski sistemi in operacije Space systems and
operations
2003-01.Slovenski inštitut za standardizacijo. Razmnoževanje celote ali delov tega standarda ni dovoljeno.

TECHNICAL REPORT
CEN/CLC/TR 17063-31-
RAPPORT TECHNIQUE
TECHNISCHER BERICHT
August 2021
ICS 49.140
English version
Space Engineering - Thermal design handbook - Part 15:
Existing Satellites
Ingénierie spatiale - Manuel de conception thermique - Raumfahrttechnik - Handbuch für thermisches Design -
Partie 15 : Véhicules spatiaux existants Teil 15: Existierende Satelliten

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 17063-31-15:2021 E
reserved worldwide for CEN national Members and for
CENELEC Members.
Table of contents
European Foreword . 9
1 Scope . 10
2 References . 11
3 Terms, definitions and symbols . 12
3.1 Terms and definitions . 12
3.2 Abbreviated terms. 12
3.3 Symbols . 17
4 International ultraviolet explorer (IUE) . 18
4.1 Mission . 18
4.2 Main subsystems . 18
4.3 Main characteristics of the satellite . 20
4.4 Orbit . 21
4.5 Thermal design requirements . 22
4.6 Design tradeoffs . 24
4.7 Thermal control of various components . 24
4.8 Estimated on orbit performance . 25
5 Orbital test satellite (OTS) . 29
5.1 Mission . 29
5.2 Main subsystems . 29
5.3 Main characteristics of the satellite . 32
5.4 Orbit . 35
5.5 Thermal design requirements . 35
5.6 Design tradeoffs . 36
5.7 Thermal control of various components . 36
5.8 Estimated on orbit performance . 42
5.9 Measured in orbit performance . 42
6 Landsat D . 49
6.1 Mission . 49
6.2 Main subsystems . 49
6.3 Main characteristics of the satellite: . 50
6.4 Orbit . 51
6.5 Thermal design requirements . 51
6.6 Design tradeoffs . 52
6.7 Thermal control of various components . 52
6.8 Estimated on orbit performance . 54
6.9 Verification . 56
6.10 Measured on orbit performance . 57
7 Infrared astronomical satellite (IRAS) . 58
7.1 Mission . 58
7.2 Main subsystems . 58
7.3 Spacecraft main characteristics . 60
7.4 Orbit . 61
7.5 Thermal design requirements . 62
7.6 Design constraints . 63
7.7 Thermal control of various components . 64
7.8 Test of the spacecraft system . 67
7.9 Test of the superfluid Helium Dewar . 68
7.9.1 General . 68
7.9.2 Test of the plug . 69
7.9.3 Prelaunch preparations . 70
7.10 On orbit performance of the spacecraft . 71
7.11 On orbit performance of the cryogenic system . 72
8 Satellite probatoire d’observation de la terre (SPOT). 76
8.1 Mission . 76
8.2 Main subsystems . 76
8.3 Main characteristics of the satellite . 77
8.4 Orbit . 80
8.5 Thermal design requirements . 80
8.5.1 Functional modes . 80
8.5.2 Orbital constraints . 80
8.5.3 Limiting temperatures . 81
8.5.4 Thermal interfaces . 83
8.6 Design tradeoffs . 83
8.7 Thermal control of various components . 84
8.7.1 Platform . 84
8.7.2 Batteries compartment . 85
8.7.3 High-resolution visible range instruments . 87
8.7.4 Payload telemetry system . 90
8.8 Estimated on-orbit performance . 92
8.8.1 Platform . 92
8.8.2 Batteries compartment . 93
8.8.3 High-resolution visible range instrument . 95
8.8.4 Payload telemetry system . 95
9 Olympus-1 . 97
9.1 Mission . 97
9.2 Main subsystems . 97
9.3 Orbit . 102
9.4 Thermal design requirements . 102
9.5 Thermal control . 102
9.6 Thermal test of olympus-1 . 105
9.6.1 Thermal vacuum test . 106
9.6.2 Infrared test . 109
10 ERS-1 . 114
10.1 Mission . 114
10.2 Main subsystems . 115
10.3 Orbit . 119
10.4 Thermal design requirements . 119
10.5 Thermal control . 122
10.6 Thermal tests . 126
10.6.1 Thermal balance test of the engineering model . 126
10.6.2 Thermal vacuum test . 132
Bibliography . 133

Figures
Figure 4-1: IUE spacecraft in orbital flight. . 18
Figure 4-2: Exploded view of the IUE spacecraft. . 20
Figure 4-3: IUE orientation to the Sun and reference axes. . 22
Figure 4-4: Assembled IUE Spacecraft. From Skladany & Seivold (1976) [42]. Notice
that this figure, which corresponds to an earlier development, differs from
Figure 4-1 in minor details. . 23
Figure 4-5: IUE main equipment platform. From Skladany & Seivold (1976) [42]. . 24
Figure 5-1: OTS mission event sequence. From Collette & Stockwell (1976) [14]. . 29
Figure 5-2: Exploded view of the OTS spacecraft. From Bouchez, Howle & Stümpel
(1978) [9]. . 33
Figure 5-3: OTS main organic diagram. From Collette & Stockwell (1976) [14]. . 34
Figure 5-4: OTS Thermal Control Subsystem temperature limits. From Stümpel (1978)a
[45]. . 35
Figure 5-5: OTS thermal control layout summary. From Stümpel (1978)a [45]. . 39
Figure 5-6: Insulation in the OTS hydrazine line system. From Stümpel (1978)a [45]. . 39
Figure 5-7: OTS heater switching diagram. . 40
Figure 5-8: Thermal insulation of the hydrazine tank. The tank is totally covered with
low emittance tape. Heaters are of the foil type (see ECSS-E-HB-31-01
Part 11, clause 4.2). The tank contacts the platform via a low conductance
amount. From Stümpel (1978)b [46]. . 40
Figure 5-9: Thermal decoupling of FCV from TCA onboard OTS. The heat barrier
maintains temperature differences up to 800 K via a length of 0,03 m. . 41
Figure 5-10: Histograms for ground and first orbit test. From Bouchez & Gülpen (1980)
[5]. The ordinates show the number of samples the temperature deviation
of which stays within the limits shown in abscissae. (∆T = T −T ). . 43
measured predicted
Figure 5-11: Histograms for orbit tests during different summer solstices. Data for 1978
and 1980 are from Bouchez & Gülpen (1981) [5] and those for 1981 from
Bouchez & Howle (1982) [7]. . 44
Figure 5-12: Temperature increases ∆T as a function of time, t elapsed since Jan 1,
1978. From Chalmers, Konzok, Bouchez & Howlw (1983) [13]. Circle:
Summer Solstice test points. Square: Winter Solstice test points. Triangle:
Equinox test points. . 46
Figure 5-13: Mean solar absorptance, α , on antenna dish white S-13 G/LO paint. From
s
Chalmers, Konzok, Bouchez & Howle (1983) [13]. Circle: Summer Solstice
test points. Square: Winter Solstice test points. Triangle: Equinox test
points. . 48
Figure 6-1: Landsat spacecraft in orbital flight. . 49
Figure 6-2: Exploded view of the Landsat D spacecraft before deployment. . 50
Figure 6-3: Assembled Wide Band Module. . 53
Figure 6-4: Thermal Control coatings used on Landsat D. . 54
Figure 7-1: IRAS spacecraft in orbital flight. See also Table 7-1. From Van Leeuwen
(1983) [53]. . 58
Figure 7-2: IRAS telescope subsystem. From Urbach et al. (1982) [52] . 61
Figure 7-3: IRAS attitude constraints during mission. From Van Leeuwen (1983) [53]. 63
Figure 7-4: IRAS spacecraft thermal control layout summary. From Van Leewen (1983,
1985) [53] & [54]. . 65
Figure 7-5: IRAS Telescope thermal control layout summary. From Urbach et al. (1982)
[52] and Sherman (1982) [41]. . 67
Figure 7-6: IRAS Test Configuration. a. Thermal model. b. Complete satellite in JPL
facility. From Van Leeuwen (1983) [53]. . 67
Figure 7-7: Effect of Critical parameters on heat load to cryogen. From Urbach,
Hopkins & Mason (1983) [50]. . 69
Figure 7-8: Tilting of the MCT for porous plug submersion. From Petrac & Mason
(1984) [39]. . 70
Figure 7-9: Vapor mass flow rate, m, and heat transfer rate, Q, through the plug vs.
pressure drop, ∆p. From Petrac & Mason (1984) [39]. 70
Figure 7-10: Histogram for ground and orbit test just after launching. The temperature
deviation is ∆T = T − T . From Van Leeuwen (1983) [53]. . 71
measured predicted
Figure 7-11: FSSS temperature, T, as a function of time, t, elapsed after launch. From
Van Leeuwen (1983) [53]. A thermal misalignment phenomenon, occurred
during the experimental phase of the mission, has been reported by
Karsten & Teule (1984) [31]. This phenomenon, which was adequately
modelled and partially overcome, was responsible for the development of
cross-scan attitude errors of up to 100 arcsec. The origins of the
misalignment changes could be traced to both spacecraft structure and
FSSS brackets. . 72
Figure 7-12: Cryogenic System Equilibrium Temperatures. From Urbach & Mason
(1984) [51]. . 74
Figure 7-13: Cryogenic boil-off rate according to different models. From Urbach,
Hopkings & Mason (1983) [50]. . 75
Figure 8-1: SPOT 1 spacecraft in orbital flight. . 76
Figure 8-2: Exploded view of the SPOT 1 subsystems and components which require
thermal control. Drawn by the compiler after Alet & Foret (1983) [1],
Fagnoni (1983) [20], Courtois & Weill (1985) [16]. Encircled numbers in the
figure are the same as those of the clauses in the text. . 84
Figure 8-3: Battery assembly of the SPOT multimission platform. From Fagnoni (1983)
[20]. . 86
Figure 8-4: Exploded view of the HRVs. From Mauduyt, Bonnet & Toulemont (1983)
[34]. . 87
Figure 8-5: Design hot mission profile for HRV and TMCU. From Racaud, d’Antin &
Lelièvret (1983) [40]. . 88
Figure 8-6: Thermal control layout summary of the HRV. From Mauduyt, Bonnet &
Toulemont (1983) [34]. . 90
Figure 8-7: SPOT 1 Satellite as seen from the –Z side. From Racaud et al. (1983) [40]. . 91
Figure 8-8: Temperature limits of the SPOT 1 platform components. From Alet & Foret
(1983) [1]. . 93
Figure 8-9: Test configuration of the batteries compartment of the SPOT multimission
platform. From Fanoni (1983) [20]. . 94
Figure 9-1: Olympus-1 in orbital flight. From Bonhomme & Steels (1984) [4], Steels &
Baston (1986) [44]. . 97
Figure 9-2: Exploded view of Olympus-1 satellite. From ESA (1984), Bowles (1987)
[10], Paul (1989) [38]. . 98
Figure 9-3: Schematic of the different phases of the Olympus-1 solar array deployment.
Prepared by the compiler after Bonhome & Steels (1984) [4], Bowles
(1987) [10]. . 100
Figure 9-4: Olympus-1 satellite thermal control layout used for thermal vacuum tests.
From Boggiatto, Colizzi, Perotto & Tavera (1985) [3]. Explanation is given
in Table 9-3. . 103
Figure 9-5: Olympus-1 satellite battery thermal control layout. a) Ni-Cd battery; b) Ni-H
battery. From Konzok, Gutschmidt, Stümpel, Schlitt & Dunbar (1987) [33]. . 105
Figure 9-6: Temperature Difference Histograms for the three test cases considered in
the Thermal Vacuum Tests of Olympus-1 satellite (see Table 9-6 above).
From Boggiatto, Colizzi, Perotto & Tavera (1985) [3]. 109
Figure 9-7: Infrared test related activities. From Messidoro & Colizzi (1986) [37]. . 111
Figure 9-8: Temperature vs. time profiles of Olympus-1 satellite as obtained from the
infrared test. North radiator, inner face. South radiator, outer
face. Communications Module – Service Module, central cylinder.
Communications Module, upper floor. From Messidoro & Colizzi
(1986) [37]. . 113
Figure 10-1: ERS-1 in flight configuration. From Francis et al. (1991) [21]. . 115
Figure 10-2: Exploded view of ERS-1 satellite. From Francis et al. (1991) [21]. . 116
Figure 10-3: Schematic of the different phases of ERS-1 SAR Antenna deployment.
From Francis et al. (1991) [21]. . 123
Figure 10-4: ERS-1 satellite. PEM external thermal design. From Haimler, Overbosch &
Pieper (1987) [24] . 124
Figure 10-5: ERS-1 satellite. PEM internal thermal design. From Haimler, Overbosch &
Pieper (1987) [24]. . 125
Figure 10-6: Temperature difference histograms for the PL-Off Phase. From Haimler,
Kamp & Pieper (1990). . 131
Figure 10-7: Transient temperature behaviour of IDHT TWT’s: a) Predicted, b)
measured. From Haimler, Kamp & Pieper (1990). . 131

Tables
Table 4-1: Characteristics of the IUE Main Subsystems . 19
Table 4-2: IUE Flight Segment Mass Summary . 21
Table 4-3: Thermal Design Requirements . 23
Table 4-4: Estimated and Measured Performance of Spacecraft Components and
Scientific Instrument Components with Nominal Power Dissipation. . 26
Table 5-1: Characteristics of the OTS main Subsystems . 30
Table 5-2: OTS Mass Summary . 33
Table 5-3: Sensor Distribution . 42
Table 5-4: In Orbit Measured Values and Curve Fitting Values . 45
Table 5-5: Change in Solar Absorptance, ∆αs, of OSR vs. Exposure Time as Deduced
from OTS Solstice Data . 47
Table 6-1: Landsat D Flight Segment Mass Summary . 51
Table 6-2: Thermal Design Requirements . 52
Table 6-3: Estimated on Orbit Performance of the Instrument Module Components . 55
Table 7-1: IRAS Main Subsystems . 59
Table 7-2: Thermal Design Requirements . 62
Table 7-3: Cryogenic System performance Summary . 72
Table 8-1: Characteristics of the SPOT 1 Main Subsystems . 77
Table 8-2: SPOT 1 Mass Summary . 79
Table 8-3: Limiting Temperatures and Heat Dissipation Rates of Typical Components –
SPOT 1 Satellite . 81
Table 8-4: Estimated and Measured Performance of the SPOT Multimission Platform
Batteries Compartment (T in K). . 94
Table 9-1: Olympus-1 Main Subsystems . 99
Table 9-2: Olympus Payload . 100
Table 9-3: Payload Subsystems Identification in Figure 9-4. . 103
Table 9-4: Olympus-1 Battery Performance Characteristics . 104
Table 9-5: Olympus-1 Thermal Test . 105
Table 9-6: Representative Cases Considered in the Thermal Test . 106
Table 9-7: Subsystem Temperature [K] after Different Steps in the Test-Mathematical
Model Interaction. . 108
-2
Table 9-8: Winter Solstice Heat Transfer Rates, Qe[W.m ], Measured and Compared
with the Requirements . 112
Table 10-1: Payload Main Subsystems . 117
Table 10-2: Typical Design Temperature Limits and PEM Dissipations . 120
Table 10-3: ERS-1 Thermal Test . 126
Table 10-4: Thermal Balance Test Phases. From Haimler, Kamp and Pieper (1990) . 128
Table 10-5: Final Level Correlation Status. Average Measured Predicted Deviation for
Steady State Case . 132

European Foreword
This document (CEN/CLC/TR 17603-31-15:2021) 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 17603-31-15:2021) originates from ECSS-E-HB-31-01 Part 15A.
Attention is drawn to the possibility that some of the elements of this document may be the subject of
patent rights. CEN [and/or CENELEC] shall not be held responsible for identifying any or all such
patent rights.
This document has been prepared under a mandate given to CEN by the European Commission and
the European Free Trade Association.
This document has been developed to cover specifically space systems and has therefore precedence
over any TR covering the same scope but with a wider domain of applicability (e.g.: aerospace).
Scope
In this Part 15, existing satellites are described and examined from a thermal control and design view.
The thermal control requirements are given and an assessment is made of the thermal control systems
used against performance for each satellite.

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

References
EN Reference Reference in text Title
EN 16601-00-01 ECSS-S-ST-00-01 ECSS System - Glossary of terms
TR 17603-31-03 ECSS-E-HB-31-01 Part 3 Thermal design handbook – Part 3:
Spacecraft Surface Temperature
TR 17603-31-05 ECSS-E-HB-31-01 Part 5 Thermal design handbook – Part 5:
Structural Materials: Metallic and
Composite
TR 17603-31-06 ECSS-E-HB-31-01 Part 6 Thermal design handbook – Part 6:
Thermal Control Surfaces
TR 17603-31-07 ECSS-E-HB-31-01 Part 7 Thermal design handbook – Part 7:
Insulations
TR 17603-31-08 ECSS-E-HB-31-01 Part 8 Thermal design handbook – Part 8: Heat
Pipes
TR 17603-31-09 ECSS-E-HB-31-01 Part 9 Thermal design handbook – Part 9:
Radiators
TR 17603-31-11 ECSS-E-HB-31-01 Part 11 Thermal design handbook – Part 11:
Electrical Heating
TR 17603-31-12 ECSS-E-HB-31-01 Part 12 Thermal design handbook – Part 12:
Louvers
TR 17603-31-13 ECSS-E-HB-31-01 Part 13 Thermal design handbook – Part 13: Fluid
Loops
TR 17603-31-14 ECSS-E-HB-31-01 Part 14 Thermal design handbook – Part 14:
Cryogenic Cooling
All other references made to publications in this Part are listed, alphabetically, in the Bibliography.
Terms, definitions and symbols
3.1 Terms and definitions
For the purpose of this Standard, the terms and definitions given in ECSS-S-ST-00-01 apply.
3.2 Abbreviated terms
The following abbreviated terms are defined and used within this Standard.
apogee boost motor
ABM
acquisition camera module
ACM
attitude control system
ACS
Clause 8: attitude control sensors
(association Francaise de normalisation), French
AFNOR
standards association.
assembly, integration and testing
AIT
active microwave instrument
AMI
attitude and orbit control system
AOCS
advanced on-board processor
AOP
along track scanning radiometer
ATSR
bearing and power transfer assembly
BAPTA
ball aerospace system division
BASD
bus coupling unit
BCU
Clause 8: burst disc
BD
Clause 9: (boitier de détection), detector housing
beginning of life
BOL
back side reflection cell
BSR
coupled charge device
CCD
constant conductance heat pipe
CCHP
control electronic unit
CEU
communication module
CM
(centre national d'etudes spatiales), French space
CNES
agency
command relay unit
CRU
coarse sun sensor
CSS
(charge utile) payload
CU
depth of discharge
d.o.d.
Dutch additional experiment
DAX
direct broadcast service
DBS
direct current
DC
direct energy transfer
DET
equinox
E
attitude measurement and control electronics
EAIM
European communication satellite
ECS
electrical ground support equipment
EGSE
equivalent isotropic radiated power
EIRP
engineering model
EM
(electronique de messure d'attitude), gyro electronics
EMA
end of life
EOL
electronic power conditioner
EPC
European remote sensing satellite
ERS
engineering test unit
ETU
fuel control valve
FCV
frequency-division multiple access
FDMA
fine error sensor
FES
fixed momentum wheel
FMW
focal plane assembly
FPA
fine sun sensor
FSSS
geostationary
GEO
global positioning system
GPS
global reference mission
GRM
ground support equipment.
GSE
goddard space flight center
GSFC
gyro sensor
GYRS
hydrazine auxiliary propulsion system.
HAPS
heat pipe
HP
high power amplifier
HPA
high resolution visible
HRV
high voltage
HV
instrument data handling and transmission system
IDTH
instrument module
IM
infrared
IR
inertial reference assembly
IRA
infrared astronomical satellite
IRAS
infrared earth sensor
IRES
infrared experiment
IRX
international ultraviolet explorer
IUE
laser retro-reflector
LRR
largest space simulation
LSS
low thrust vents
LTV
low voltage
LV
mission adapter module
MAM
magnetic coil
MCL
maritime communication satellite
MCS
main cryogenic tank
MCT
(mécanisme d'entrainement du générateur solaire)
MEGS
solar array driver system
multilayer insulation.
MLI
multimission modular spacecraft
MMS
(module de propulsion) propulsion module
MP
(module de servitude) service module
MS
multispectral scanner
MSS
(modèle structural et thermique), structural and
MSTH
thermal model
on board computer
OBC
overall check-out equipment
OCOE
optical solar reflector
OSR
orbital test satellite
OTS
panchromatic
Pan
power control unit
PCU
power distribution unit
PDU
payload electronics module
PEM
platform
PF
plateforme multimission
PFM
payload
PL
porous plug
PP
(plateau porte equipements), equipment base plate
PPE
precision range and range rate equipment
PRARE
programmable power supply
PROPOS
power supply electronics
PSE
radar altimeter
RA
reaction control system
RCS
repeater drive unit
RDU
radio frequency
RF
rate integrating gyro
RIG
remote interface unit/expander unit
RIU/EU
solstice
S
synthetic aperture radiation
SAR
spacecraft component on telescope
SCOT
solar environment simulator
SES
super high frequency
SHF
service module
SM
(satellite probatoire d'observation de la Terre), trial
SPOT
Earth observation satellite (in the beginning)
(satellite Pour l'observation de la Terre), Earth
observation satellite (current usage) (Dyson (1986)
[18])
summer solstice
SS
second surface mirror
SSM
satellite-switched time-division multiple access
SS-TDMA
(senseur terrestre digital), digital Earth sensor
STD
thrust chamber assembly
TCA
thermal control subsystem
TCS
test data handling
TDH
tracking and data relay satellite
TDRS
thermal interface mathematical model
TIMM
thematic mapper
TM
(télémesure de la charge utile), payload telemetry
TMCU
system
telemetry, tracking and command
TT&C
tape unit
TU
television broadcast system
TVBS
travelling wave tube
TWT
travelling wave tube amplifier
TWTA
ultra-violet
UV
vapor cooled shield
VCS
very high frequency
VHF
wide band module
WBM
multispectral
XS
3.3 Symbols
heat transfer rate, [W]
Q
temperature, [K]
T
characteristic temperature, [K]
Tm
effective heat pipe length, [m]
leff
mass, [kg]
m
−
louver heat rejection capability, [W.m ]
q
time, [d]
t
reference time, [d]
t0
solar absorptance
αs
attitude angle of the satellite with respect to Sun line,
β
[angular degrees] it is not defined in a unified way
Clauses 9 and 12: angle between the line to the Sun
and the longitudinal axis of the satellite
Clause 10: angle between the lie to the Sun and the
normal to orbit plane
Clause 11: 90° minus the angle between the line to the
Sun and the normal to orbit plane
hemispherical total emittance
ε
standard deviation
σ
τ characteristic time
International ultraviolet explorer (IUE)
4.1 Mission
−
Observation of the ultraviolet spectra of astronomical sources in the spectral region between 1,15x10
−
m and 3,2x10 m.
Launching date: Jan. 26, 1978.
Mission life: 3 years design life with a 5 years design goal. After six year in orbit, a detailed study by
NASA shown that the solar panels would support the operation till 1988.

Figure 4-1: IUE spacecraft in orbital flight.
4.2 Main subsystems
IUE carries a 0,45 m diameter Casegrain telescope, which uses an echelle spectrograph for UV
− − −
7 7 11
astronomy in the spectral region between 1,15x10 m and 3,2x10 m. Aimed resolution was 10 m in
th
the mentioned region of the spectrum for stars and planets brighter than 7 visual magnitude, and
−
11 th
lower resolution (6x10 m) for stellar and extended objects as faint as 12 magnitude. Spectroscopy
th
on stars as faint as 18 magnitude has been performed with this instrument.
Relevant characteristics of the different subsystems are summarized in Table 4-1.
Table 4-1: Characteristics of the IUE Main Subsystems
Subsystem Purpose Components
Power Providing regulated (28 V
• Two deployable solar panels furnished by
Direct Energy DC ± 2%) power which is
ESA. Manufactured by Aérospatiale
transfer (DET) transferred from the solar
(France) and AEG-Telefunken
array to the spacecraft bus.
(Germany), see Bulloch (1978) [11].
• Power supply electronics (PSE) composed
of redundant power modules. A power
module consists of battery charger, shunt
driver, boost regulator, control unit and
12 dump resistors.
• Mission Adapter Module (MAM)
interfacing modules to one another, to
solar arrays and to batteries.
• Two 21,6x10 C (3,6 A.h) Ni-Cd batteries.
Communication Transmitting data to
• VHF System used during transfer orbit
ground. Receiving ground-
and for tracking during mission orbit. It
generated commands.
consists of two redundant transponders,
Providing range and range
antenna distribution and a turnstile
rate signals for metric
antenna system.
tracking.
• S band downlink system for mission
orbit. Two redundant transmitters and
four S-band antennae. Active antenna is
selectable on ground.
Commend & Accepting and decoding
• Two redundant command decoders
Data Handling commands from ground or
processing messages either from VHF
from onboard computer.
receiver analog signal or from digital
Encoding spacecraft and
information generated by the onboard
scientific instrument
computer.
telemetry. Performing
• A Command Relay Unit (CRU).
attitude control
computations. Monitoring • Two data multiplex units. Each one
critical subsystems and
consists of dataplexer, analog subplexer
controlling exposure times and digital subplexer.
of spectral images.
• Advanced On-board Processor (AOP)
computer.
Stabilization & Precessing the spacecraft
• Earth and Sun sensors for ground
Control spin axis 180° in
computer attitude determination.
preparation for the apogee
• Rate gyros and analog Sun sensors for
engine burn.
initial spacecraft acquisition.
Attaining the correct orbit.
• Inertial Reference Assembly (IRA). Six
Despinning once the orbit
gas-bearing, pulse rebalanced, rate-
has been achieved.
integrating gyros.
Subsystem Purpose Components
Controlled pointing and
• Redundant two-axis digital Sun sensor
slew of the telescope
system.
according to specifications.
• Redundant set of star trackers within the
scientific instruments and utilizing the
telescope optics.
• Redundant reacting wheel.
• Nutation reaction accelerometers.
• Apogee Boost Motor (ABM) and
Hydrazine Auxiliary Propulsion System
(HAPS)
4.3 Main characteristics of the satellite
The main body is octogonal in shape. The telescope extends from the upper end of the main body. It is
attached to the spacecraft structure by means of s strong ring resting on three columns which carry the
load to the lower spacecraft structure. The columns are supported laterally by truss members of the
main body.
Figure 4-2: Exploded view of the IUE spacecraft.
...