GAL-REQ-ASTD-SA-R-0003 3.2 2006-01-16emits.sso.esa.int/emits-doc/ESTEC/AO-1-5524-RD1_GAL-REQ... · 2010. 1. 13. · Fischer, Horst-Dieter Sauerer, Bernhard Franz, Horst-Jürgen X

Galileo IOV – Space Segment Environment Requiremen ts and Test Specification The contents of this document are the copyright of EADS Astrium GmbH. All rights reserved.

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Doc-No.: GAL-REQ-ASTD-SA-R-0003 Issue: 3.3 Date: 16 January 2006 Page: 2 of 163 UNCLASSIFIED

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Baker, Roger X Mayer, Rupert X Berberich, Dr. Stefan X McCrorie, Claire Böck, Anna-Maria Mediavilla, Ignacio Bögel, Gerold X Meyer, Alexandra Cossu, Cinzia Neundorf, Jürgen Cox, Alan Nietner, Gerhard X Desroches, Catherine X Paus, Stefan Dietz, Gerhard X Porte, Dr. François Ebert, Dr. Klaus Reichel, Jan Elsner, Günther Reinhold, Norbert Feragalli, Roberto Sasse, Franz Fischer, Horst-Dieter Sauerer, Bernhard Franz, Horst-Jürgen X Schachenmeier, Markus X Freidl, Erwin X Schneider, Matthias Furones, Valentin Schön, Georg Goeldner, Torsten X Schuster, Ludwig Hamon, Stéphane Smargiassi, Michele Hecking, Peter Smith, David X Hendricks, Reinhard X Spatafora, Vincenzo Hertle, Anton Squillaci, Jean-Régis X Honold, Hans-Peter Tomi, Hervé Hörl, Dr. Kay-Uwe Tucker, Paul X Janssen, Greetje Vogel, Matthias X Jenkins, Ian Robert Wachter, Claudia Kaufmann, Christiane Weichs, Erich X Keding, Hans-Jürgen Widmann, Hans King, Raymond Wolframm, Aribert Klein, Miriam Ziegler, Franz-Josef Lindenthal, Werner X Luc, André Luck, Stephen Maggi, Emanuele X Marco, Victor X




Document Change Record

Issue / Rev. Date Changes

3/1 27/06/2005 Update taking into account the results of the technical discussions with the subsystem contractors.

Chapter 2.3 Update of the acronyms table

Chapter 3.4 change of the satellite acceptance sine test factor (1.0 or flight level is required).

Chapter 3.4 definition of the unit test factor requirement.

Chapter 4.1.1.1.1 clarification of the levels applicable to the containers.

Chapter 4.1.1.1.2 definition of the ground loads during launch site operations.

Chapter 4.1.1.2.2 correction of errors on the satellite sine level for multiple launch.

Chapter 4.1.2.1.3 suppression of a misleading precision regarding storage requirements.

Chapter 4.1.2.3.1 addition of the maximum angular rates applicable in orbit.

Chapter 4.2.2.3 modification of the platform unit’s configuration: definition of ES configuration and pressure transducer off during launch up to separation.

Chapter 4.2.2.4 update of the payload unit’s configuration (ion pumps off at launch)

Chapter 4.2.6.3 correction of referred section.

Chapter 4.2.7 correction of referred section.

Chapter 4.2.15.1.3 start-up temperature not applicable to SAHD.

Chapter 4.2.15.1.4 Separate tables for CSS and FSS

Chapter 4.2.15.1.4 IRES renamed ES

Chapter 4.2.15.1.4 RWL renamed RW

Chapter 4.2.15.1.4 Modification of the ICDU start-up temperature

Chapter 4.2.15.1.4 Modification of the ES start-up temperature

Chapter 4.2.15.1.4 Modification of the gyro start-up temperature

Chapter 4.2.15.1.4 Modification of the MTR start-up temperature





Chapter 4.2.15.1.4 Modification of the RW start-up temperature

Chapter 4.2.15.1.5 Modification of the thrusters temperatures after clarification by the propulsion contractor.

Chapter 4.2.15.2 suppression of the Power and Data Handling harness temperature requirements form the payload section (not applicable to payload)

Chapter 4.2.15.2 Modification of the FGUU start-up temperature

Chapter 4.2.15.2 Modification of the MISREC start-up temperature

Chapter 4.2.15.2 Modification of the MISSP start-up temperature

Chapter 4.2.15.2 Modification of the PLSU temperatures

Chapter 4.2.15.2 Modification of the NSGU temperatures

Chapter 4.2.15.2 Modification of the FRTC temperatures

Chapter 4.2.15.2 Modification of the CRTC temperatures

Chapter 4.2.15.2 Modification of the BBTC temperatures

Chapter 4.2.15.2 Modification of the IPTC temperatures

Chapter 4.2.15.2 Definition of the SAR harness temperatures

Chapter 4.2.15.2 Modification of the splitter temperatures

Chapter 4.2.15.2 Modification of the HP switch temperatures

Chapter 4.2.15.2 Modification of the HP load temperatures

Chapter 4.2.15.2 Modification of the OPTC temperatures

Chapter 4.2.15.2 Modification of the MISTC temperatures

Chapter 4.2.15.2 Modification of the SARIPTC temperatures

Chapter 4.2.15.2 Modification of the SAROPTC temperatures

Chapter 4.2.15.2 Modification of the HP harness temperatures

Chapter 4.2.15.4 Introduction of heatpipe temperatures

Chapter 4.2.15.5 Power and Data Handling harness temperature requirements placed on a dedicated section.

Chapter 4.2.16 Introduction of a new section defining the interface temperatures and the satellite external thermo-optical properties applicable to the external units.

Chapter 6 addition of a requirement on the satellite magnetic field environment.





Chapter 8.1.1.1 Suppression of the reference to section AD.

Chapter 8.1.3 indicated deleted.

Chapter 8.1.4.1, 8.1.4.4, 8.1.4.5 MRB replaced by NRB

Chapter 8.1.5 improvement of wording

Chapter 8.1.6 addition of the term ”resonance search” for clarification

Chapter 8.1.6.1 Addition of a requirement to clarify the mechanical environmental requirements for units which include support brackets.

Chapter 8.1.6.1 correction of the references to the micro-vibration environment section.

Chapter 8.1.6.1 Definition of the applicable mechanical environment (quasi-static, sine, random and shock) for each unit.

Chapter 8.1.6.2 Addition of a specific quasi-static level for the propulsion tank

Chapter 8.1.6.3 Addition of a specific sine level for the propulsion tank

Chapter 8.1.6.4 Correction of a typo on the in-plane random tables (frequency value)

Chapter 8.1.6.4 Clarification of the random level applicable zones

Chapter 8.1.6.5 Clarification of the acoustic test durations

Chapter 8.1.6.6 Relaxation of the shock test levels taking into account more precisely the definition of the satellite structure

Chapter 8.1.7.1.5 Correction of the number of acceptance cycles

Chapter 8.1.7.1.6 Correction of the qualification thermal test sequence starting with a hot cycle to achieve unit out-gassing

Chapter 8.1.7.1.6 Modification of the temperature stabilisation criteria

Chapter 8.1.7.1.6 Update of the dwell duration

Chapter 8.1.7.2.2 Correction of the acceptance thermal test sequence starting with a hot cycle to achieve unit out-gassing

Chapter 8.1.7.2.2 Modification of the temperature stabilisation criteria





3.2 05.12.2005 Specification transferred to DOORS. Requirements content identical to Issue 3.1 (except the changes which are described hereafter) but all requirements now have a unique requirement id number

Chapter 1 Introduction of the wording providing the rules to use the pdf and Doors module files

3.3 16.01.2006 §4.2.2.2 Figure 6 'Satellite Phases' updated §4.2.2.2 ENVREQ-211 Figure 7 'Satellite modes' updated

§4.2.2.3 ENVREQ-215

Name of ‘Test Mode’ changed to name ‘Stand By/Test Mode.

GYRO 2 deleted, There is now only one Gyro The Gyro status in Intermediate Safe Mode changed from old: T-op (T-operating)

new: Tnon-op (Tnon-operating)

SS1, SS2, SS3 with internal redundancy changed to FSS1 (one single unit), FSS2 (one single unit) and CSS1 (one single unit) and CSS2(one single unit),

SADM’s are on in sun acquisition mode

Magnetic Torquer are on in intermediate safe mode

Propulsion thrusters are off in intermediate safe mode

§4.2.2.4 ENVREQ-220

Name of ‘Test Mode’ changed to name ‘Stand By/Test Mode.

FRTC1-2, CRTC1-2 and BBTC1-6 deleted

OPF1 = Output Filter 1included

OPF2 = Output Filter 2 included

MISREC 2 deleted, There is now only one MISREC

MISSP and MISTC deleted

The status of RAF1 changed in EAM: old: Tnon-op (Tnon-operating) new: T-op (T-operating)




Issue / Rev. Date Changes The status of PHM1 changed in EAM: old: Tnon-op (Tnon-operating) new: T-op (T-operating) The status of CMCU changed in EAM: old: Tnon-op (Tnon-operating) new: T-op (T-operating) The status of RTU changed in EAM: old: Tnon-op (Tnon-operating) new: T-op (T-operating)

§4.2.15.2 ENVREQ-486 The number of NAVHP increased in the list to 7.

§4.2.15.1.4 ENVREQ-406

The label of the fine sun sensor changed from SS1 to FSS1 and in addition the FSS2 included

§4.2.15.1.4 ENVREQ-409

The label of the coarse sun sensor changed form SS2 & SS3 to CSS1 & CSS2. CSS temperature range updated according to unit characteristics

§4.2.15.1.4 ENVREQ-415

The GYRO 2 deleted

§4.2.15.2 ENVREQ-507

The MISREC2 deleted §4.2.15.2 ENVREQ-510 The requirement deleted, because MISSP deleted §4.2.15.2 ENVREQ-513 The requirement deleted, because MISTC deleted §4.2.15.2 ENVREQ-471 The requirement deleted, because FRTC1 & 2 deleted §4.2.15.2 ENVREQ-474 The requirement deleted, because CRTC1 & 2 deleted §4.2.15.2 ENVREQ-477 The requirement deleted, because BBTC1 & 2 deleted




Issue / Rev. Date Changes §8.1.7.1.1 ENVREQ-945 New requirement included: For internally redundant units, the unit supplier shall perform performance qualification test for the prime and redundant parts, with switching from one to the other. §8.1.7.2.1 ENVREQ-946 For internally redundant units, the unit supplier shall perform performance acceptance test for the prime and redundant parts, with switching from one to the other. §4.2.1.5 ENVREQ-195

Wording changed: Old wording: Specifically for the clocks, the temperature margin between the predicted temperature and the design temperature shall then be equal to 8°C in hot case and 3 °C in cold case. Therefore clocks will be nominally regulated at 0 °C. The thermal design margin for the clocks shall be taken into account by the heater budget. New wording: The clocks will be nominally regulated at 5 °C for the normal orbit case. An uncertainty margin of 2°C shall be c onsidered. Furthermore the thermal design shall provide a minimum margin of 15% between maximum needed heater power and effectively installed heater power.

§4.2.15.2 ENVREQ-453

The temperature stability changed from ‘one orbit’ to 24 hours.

§4.2.15.2 ENVREQ-456


§4.2.15.2 ENVREQ-459


§4.2.15.2 ENVREQ-465


§4.2.15.2 ENVREQ-468


§4.2.15.2 ENVREQ-486


§4.2.15.2 ENVREQ-495


§4.2.15.2 ENVREQ-498






§4.2.15 ENVREQ-948

New requirement included: For the unit reliability calculation the reliability temperatures which shall be used are presented in the tables of § 4.2.15.1 (T rel.). Furthermore, Trel has been defined in the tables of § 4.2.15.1 for all relevant units. §4.2.15.3 ENVREQ-547 old status: all temperatures were TBD new status: temperatures included with TBC §4.2.15.1.3 ENVREQ-390 hot temperature limits updated: old values: qual=100°C, acc=95°C, TCS design=90°C new values: qual=130°C, acc=125°C, TCS design=120°C §4.2.8.2 Old status: Fairing jettisoning about TBD sec after lift-off New status: Fairing jettisoning about 320 sec after lift-off §8.2.1.5 ENVREQ-865 The (TBD) changed to (TBD in the TV test specification). §4.2.6.2.1 ENVREQ-262 conductive conductance changed old value: 1 W/K (TBC). new value: 0,25 W/K (TBC).

§3.4: ENVREQ-44

Change of FoSY (4.0� 2.0) and FoSU (5.0�3.0) for satellite MGSE interfaces

§4.1.1.2.5: ENVREQ 127

Deletion of the SA release shock due to the availability of a low shock system (thermal knife)

§4.2.8.2: For ZENIT the versions with the Fregat, Fregat SB and DM-SL have been deleted.

§8.1.6.4, ENVREQ-967, 970, 707, 709, 713, 715, 720, 722, 727, 730, 734 changed based on the results of the MRR#1 Vibro-acoustic analysis. ENVREQ-973, 974, 975, 976, 977, 978 included

ENVREQ-666 updated for PHM, Gyro and for RW





§4.2.2.2, Satellite Phases and Satellite Modes updated

§8.1.2, ENVREQ-606: note 1) completed by inclusion of “measurement of the inrush current”.

ENVREQ-666 random zones updated for PHM (was 699/702, is 967/970), GYRO (was 727/730, is 734) and RW (was 727/730, is 974/977)

ENVREQ-750, shock levels decreased as a result of the low shock levels by the SA release system.

Old change log (before issue 3.1) not any longer part of this document.

§3.1, table 1: launch configuration: a TBC for the accommodation on Ariane and Proton included. For Soyuz (Zenit) the definition of the single launch adapter précised.




Table of Contents

1 Introduction....................................... ............................................................ 18

2 Documents.......................................... .......................................................... 19 2.1 Applicable Documents .................................................................................... 19 2.2 Reference Documents .................................................................................... 19 2.3 Acronyms ....................................................................................................... 19

3 Definitions........................................ ............................................................. 22 3.1 Launchers....................................................................................................... 22 3.2 Coordinate Systems ....................................................................................... 24 3.2.1 S/C Coordinate System .................................................................................. 24 3.2.2 Launcher Coordinate System ......................................................................... 24 3.3 Strength.......................................................................................................... 25 3.4 Safety Factors ................................................................................................ 27

4 Environment........................................ .......................................................... 30 4.1 Mechanical Environment ................................................................................ 30 4.1.1 Satellite Level Mechanical Environment.......................................................... 30 4.1.2 Unit Level Mechanical Environment................................................................ 38 4.2 Thermal Environment ..................................................................................... 42 4.2.1 General Requirements.................................................................................... 42 4.2.2 Mission Overview ........................................................................................... 46 4.2.3 Ground Facilities Thermal Environments ........................................................ 55 4.2.4 Pre-Launch Phase.......................................................................................... 56 4.2.5 Activities in S/C Preparation Building.............................................................. 56 4.2.6 Spacecraft Configuration at Launch................................................................ 57 4.2.7 Under Fairing Conditions before Launch......................................................... 60 4.2.8 Time Line from Launch to SA Deployment...................................................... 61 4.2.9 Depressurisation under fairing ........................................................................ 63 4.2.10 Thermal conditions under fairing..................................................................... 65 4.2.11 Orbit case: 'Safe Mode '.................................................................................. 77 4.2.12 Orbit case: 'Intermediate Safe Mode ' ............................................................. 78 4.2.13 Thermal Space environment........................................................................... 80 4.2.14 Thermal in orbit environment .......................................................................... 83 4.2.15 Unit Temperature Limits ................................................................................. 84 4.2.16 Spacecraft Thermal Interfaces for External Units.......................................... 116

5 Cleanliness ........................................ ......................................................... 118

6 Electromagnetic Environment ........................ ........................................... 119

7 Orbital Environment ................................ ................................................... 120

8 Qualification and Acceptance Environmental Tests ... ............................. 121 8.1 Unit Level Tests............................................................................................ 121 8.1.1 Definition and Objective................................................................................ 121




8.1.2 Test Sequence ............................................................................................. 122 8.1.3 Deleted......................................................................................................... 123 8.1.4 General test conditions ................................................................................. 123 8.1.5 Performance checks between and during tests............................................. 124 8.1.6 Mechanical Tests.......................................................................................... 125 8.1.7 Thermal Tests............................................................................................... 138 8.1.8 Proof Pressure Test...................................................................................... 149 8.1.9 Leakage Test................................................................................................ 150 8.1.10 Other Tests................................................................................................... 150 8.2 Satellite Level Tests...................................................................................... 153 8.2.1 Structural Thermal Model (STM) Tests ......................................................... 153 8.2.2 Structural Model (SM) Tests ......................................................................... 157 8.2.3 Proto-flight Model Satellite Test .................................................................... 157 8.2.4 Spacecraft Acceptance Test Program........................................................... 162




List of Figures Figure 1 S/C Coordinate System ......................................................................................... 24 Figure 2 Launcher Coordinate System................................................................................. 25 Figure 3 Derivation of the dimensioning loads...................................................................... 26 Figure 4 Launcher separation shock spectrum (flight level).................................................. 38 Figure 5 Definition of terms applicable for thermal design.................................................... 44 Figure 6 Satellite Phases ..................................................................................................... 47 Figure 7 Satellite modes ...................................................................................................... 49 Figure 8 Soyuz and Zenit launch configuration (single and dual launch).............................. 57 Figure 9 Ariane 5 launch configuration................................................................................. 58 Figure 10 Proton launch configuration ................................................................................. 59 Figure 11 - off phase venting profile..................................................................................... 64 Figure 12 Thermal conditions under fairing .......................................................................... 65 Figure 13 Aerodynamic heating after fairing jettison ............................................................ 66 Figure 14 Coast Phase Hot Case for multiple launch (Angle between Orbiter /Spacecraft roll

axis and Sun = 60°) ................................ ...................................................................... 68 Figure 15 Coast Phase Hot Case for single launch (Angle between Orbiter /Spacecraft roll

axis and Sun = 60°) ................................ ...................................................................... 70 Figure 16 Coast Phase Hot Case for single launch (Angle between Orbiter /Spacecraft roll

axis and Sun = 0°) ................................. ....................................................................... 71 Figure 17 Coast Phase Hot Case for single launch (Angle between Orbiter /Spacecraft roll

axis and Sun = 90°) ................................ ...................................................................... 72 Figure 18 Sun Acquisition Mode conditions (Before Solar Generator deployment) .............. 73 Figure 19 Sun Acquisition Mode conditions (After Solar Generator deployment) ................. 73 Figure 20 Yaw steering law.................................................................................................. 79 Figure 21 Total solar irradiance ........................................................................................... 81 Figure 22 - angle elevation of the orbit................................................................................. 84 Figure 23 - isothermal radiative units ................................................................................ 141 Figure 24 Test arrangements for internally mounted 'conductive' units .............................. 142 Figure 25 Qualification Thermal Test Sequence ................................................................ 145 Figure 26 Acceptance Thermal Test Sequence ................................................................. 149

List of Tables Table 1 - Launchers........................................................................................................... 23 Table 2 - Safety factors...................................................................................................... 27 Table 3 - Test factors for satellite structure and integrated satellite.................................... 28 Table 4 - Factors of Safety for units ................................................................................... 28 Table 5 - Test factors for units ........................................................................................... 29 Table 6 - Transportation Containers Limit Loads................................................................ 30 Table 7 - Satellite Transportation Limit Loads .................................................................... 31 Table 8 - Quasi-static Flight Limit Loads (envelope for multiple launch configuration)........ 34 Table 9 - Quasi-static Flight Limit Loads (envelope for single launch configuration)........... 35 Table 10 - Sine longitudinal envelope (flight level) ............................................................. 35 Table 11 - Sine lateral envelope (flight level)...................................................................... 36 Table 12 - Sine longitudinal envelope (flight level) ............................................................. 36




Table 13 - Sine lateral envelope (flight level)...................................................................... 36 Table 14 - Flight Level acoustic noise spectrum................................................................. 37 Table 15 - Launcher separation induced shock.................................................................. 37 Table 16 - Handling loads .................................................................................................. 39 Table 17 - Transportation loads ......................................................................................... 39 Table 18 - Correlation of the main satellite phases and its applicable S/C modes.............. 48 Table 19 - Platform unit's configuration .............................................................................. 51 Table 20 - Payload unit's configuration .............................................................................. 53 Table 21 - Satellite mode transitions .................................................................................. 54 Table 22 - Climatic conditions on ground ........................................................................... 55 Table 23 - Transportation equipment environment............................................................. 55 Table 24 - Environment conditions in the integration/ preparation facilities and transport

containers..................................................................................................................... 56 Table 25 - Under fairing conditions before launch .............................................................. 61 Table 26 - Solar spectral irradiance ................................................................................... 82 Table 27 - Temperature limits of TX/RX (S-Band Transponder) ......................................... 85 Table 28 - Temperature limits of TX/RX (TTCHYB (Hybrid)) .............................................. 85 Table 29 - Temperature limits of TTCANTN and TTCANTZ (S-Band Antenna Nadir/Zenith)

..................................................................................................................................... 86 Table 30 - Temperature limits of TTCHAR (TTC RF-Harness)........................................... 86 Table 31 - Temperature limits of LRR (Laser Retro Reflector) ........................................... 87 Table 32 - Temperature limits of LRR (Laser Retro Reflector) ........................................... 87 Table 33 - Temperature limits of SA-Y & SA+Y (Solar Array -Y & +Y) ............................... 88 Table 34 - Temperature limits of SADM-Y & SADM+Y (Solar Array Drive Mechanism -Y &

+Y)................................................................................................................................ 89 Table 35 - Temperature limits of SAHD-Y1,2,3,4 & SAHD+Y1,2,3,4 (Solar Array Hold-down

-Y & +Y)........................................................................................................................ 89 Table 36 - Temperature limits of BATT (Battery)................................................................ 90 Table 37 - Temperature limits of ICDU (Integrated control and data handling unit) ............ 90 Table 38 - Temperature limits of FSS1 & FSS2 (Fine Sun Sensor)................................... 91 Table 39 - Temperature limits of CSS1 & CSS2 (Coarse Sun Sensors)............................ 91 Table 40 - Temperature limits of ES1 & ES2 (Earth Sensor 1 & 2)..................................... 92 Table 41 - Temperature limits of GYRO (Gyro) .................................................................. 92 Table 42 - Temperature limits of MTR1 & MTR2 (Magnetic Torquer 1 & 2)........................ 93 Table 43 - Temperature limits of RW1 & RW2 & RW3 & RW4 (Reaction Wheel 1, 2, 3, 4) 93 Table 44 - Temperature limits of PROTANK (Tank) ........................................................... 94 Table 45 - Temperature limits of PROFVV (Fill and Vent Valve) ........................................ 94 Table 46 - Temperature limits of PROPT (Pressure Transducer) ....................................... 95 Table 47 - Temperature limits of PROLF (Filter) ................................................................ 95 Table 48 - Temperature limits of PROFDV (Fill and Drain Valve)....................................... 96 Table 49 - Temperature limits of PROLVA & PROLVB (Latching Valve 1 & 2)................... 96 Table 50 - Temperature limits of PROTPA & PROTPB (Test Port A&B) ............................ 97 Table 51 - Temperature limits of PIPS (Piping) .................................................................. 97 Table 52 - Temperature limits of PRORCT1A & PRORCT2A & PRORCT3A & PRORCT4A

& PRORCT1B & PRORCT2B & PRORCT1B & PRORCT1B & (Reaction Control Thruster 1A,2A,3A,4A, 1B,2B,3B,4B)............................................................................ 98

Table 53 - Temperature limits of RAFS1 & RAFS2 (Rubidium Atomic Frequency Standard 1 & 2)............................................................................................................................... 99

Table 54 - Temperature limits of PHM1 & PHM2 (Passive Hydrogen Maser 1 & 2) ........... 99




Table 55 - Temperature limits of CMCU (Clock Monitoring and Control Unit)................... 100 Table 56 - Temperature limits of PLSU (Payload Security Unit) ....................................... 100 Table 57 - Temperature limits of PLSU (Payload Security Unit) ....................................... 101 Table 58 - Temperature limits of FGUU (Frequency Generation and Up-conversion Unit)101 Table 59 - Temperature limits of IPTC1 & IPTC2 & IPTC3 (Input Test Coupler) .............. 102 Table 60 - Temperature limits of SPLIT1 & SPLIT2 & SPLIT3 (Low Power Splitter)......... 102 Table 61 - Temperature limits of NAVHPA1 & NAVHPA2 & NAVHPA3 & NAVHPA4 &

NAVHPA5 & NAVHPA6 & NAVHPA7 (Solid State Power Amplifier) ........................... 103 Table 62 - Temperature limits of NAVSW1 & NAVSW2 & NAVSW3 (RF High Power Switch)

................................................................................................................................... 103 Table 63 - Temperature limits of NAVLOAD1 & NAVLOAD2 & NAVLOAD3 (RF High Power

Load) .......................................................................................................................... 104 Table 64 - Temperature limits of OPF1 & OPF2 (Output Filters) ..................................... 104 Table 65 - Temperature limits of OMUX (OMUX Diplexer LB (E5/E6))............................. 105 Table 66 - Temperature limits of OPTC1 & OPTC2 & OPTC3 (Output Test Couplers)..... 105 Table 67 - Temperature limits of NAVANT (L-Band Navigation Antenna)......................... 106 Table 68 - Temperature limits of MISREC (C-Band Receiver) ........................................ 106 Table 69 - Temperature limits of MISANT (C-band Mission Antenna) .............................. 107 Table 70 - Temperature limits of SARANT (SAR Rx/Tx Antenna) .................................... 107 Table 71 - Temperature limits of SARIPTC (SAR Test Coupler 406 MHz) ....................... 108 Table 72 - Temperature limits of SAROPTC (SAR Test Coupler L-Band) ........................ 108 Table 73 - Temperature limits of SART (SAR Transponder Assembly) ............................ 109 Table 74 - Temperature limits of RTU (Remote Terminal Unit) ........................................ 109 Table 75 - Temperature limits of HPHAR01 till HPHAR017 (RF Cables).......................... 110 Table 76 - Temperature limits of LPHAR (RF LP Harness) .............................................. 110 Table 77 - Temperature limits of CBHAR (C-Band Harness) and SARHAR (SAR Harness)

................................................................................................................................... 111 Table 78 - Temperature limits of STPF-x, STPF+y, STPF-y, STPF-z, STPFM, STPFSF1,

STPFSF2, STPL+x, STPL+y, STPL-y, STPL+z, S/C Structure Panels ....................... 111 Table 79 - Temperature limits of LSSIF1, LSSIF2, LSSIF3, LSSIF4, Separation Interface

................................................................................................................................... 112 Table 80 - Temperature limits of External MLI.................................................................. 112 Table 81 - Temperature limits of External High Temperature MLI .................................... 113 Table 82 - Temperature limits of Internal MLI................................................................... 113 Table 83 - Temperature limits of OSR (Optical Surface Radiator) .................................... 114 Table 84 - Temperature limits of Heaters (Foil Heaters)................................................... 114 Table 85 - Temperature limits of Thermistors................................................................... 115 Table 86 - Temperature limits of Heatpipes ..................................................................... 115 Table 87 - Temperature limits of PDHAR (Power and Data Handling Harness) ............... 116 Table 88 - Interface temperatures for external units ......................................................... 116 Table 89 - Thermo optical properties for external units .................................................... 117 Table 90 - Cleanliness Conditions.................................................................................... 118 Table 91 - Nominal Unit Test Sequence........................................................................... 122 Table 92 - Quasi-static, sine and random levels............................................................... 128 Table 93 - Sine qualification levels................................................................................... 129 Table 94 - Sun Sensors sine qualification levels .............................................................. 130 Table 95 - TTC antennas sine qualification levels ............................................................ 130 Table 96 - Propellant tank sine qualification levels (longitudinal)...................................... 130 Table 97 - Propellant tank sine qualification levels (lateral) .............................................. 130




Table 98 - Out of plane random vibration levels for units located on the +X panel 1) RAFS................................................................................................................................... 131

Table 99 - In plane random vibration levels for units located on the +X panel 1) RAFS.... 131 Table 100 - Out of plane random vibration levels for units located on the +X panels 2) PHM

................................................................................................................................... 132 Table 101 - In plane random vibration levels for units located on the +X panels 2) PHM . 132 Table 102 - Out of plane random vibration levels for units located on the -X panels

(propulsion units except tank) ..................................................................................... 132 Table 103 - In plane random vibration levels for units located on the -X panels (propulsion

units except tank)........................................................................................................ 133 Table 104 - Out of plane random vibration levels for units located on the +/- Y payload

panels......................................................................................................................... 133 Table 105 - In plane random vibration levels for units located on the +/- Y payload panels

................................................................................................................................... 133 Table 106 - Out of plane random vibration levels for units located on the +Z panel.......... 134 Table 107 - In plane random vibration levels for units located on the +Z panel ................ 134 Table 108 - Out of plane random vibration levels for units located on the platform +Y/-Y/

panels and shear frames ............................................................................................ 134 Table 109 - In plane random vibration levels for units located on the platform +Y/-Y/ panels

and shear frames........................................................................................................ 135 Table 110 - Out of plane random vibration levels for units located on the platform M panel

................................................................................................................................... 135 Table 111 - In plane random vibration levels for units located on the platform M panel.... 135 Table 112 - All three axes random vibration levels for the sun sensors and the TTC

antennas..................................................................................................................... 136 Table 113 - Applicable test tolerances ............................................................................. 159




1 Introduction This specification is one of the Galileo S/C support specifications which together with the particular satellite specifications establish the set of applicable requirements for the S/C and for the S/C units.

This specification defines: • The environment to which the spacecraft will be submitted from integration until the

end of its operational life, • The test environment to which the spacecraft shall be submitted to demonstrate its

capability to withstand relevant environmental conditions without damage, • The environment to which the spacecraft units will be submitted from integration until

the end of its operational life, • The test environment to which the spacecraft units shall be submitted to demonstrate

their capability to withstand relevant environmental conditions without damage, • The environment during transportation and storage, • The environmental conditions cover:

• mechanical environment • thermal environment • climatic environment

The EMC and the Space Environment requirements are respectively specified in AD01 and AD02. This specification defines the global spacecraft environment and contains overall requirements and guidelines for the subsystems, units and components. The details of lower level tests (e.g. subsystem, units and components tests) shall be defined in the individual test procedures and reports. This requirements specification will be delivered in the form of a Doors module and as a signed PDF file. Should differences between these two forms be detected, then they shall be brought to the attention of the Space Segment Prime for clarification. However, in general, the following rules shall apply: • if a requirement is present in one form of the document and not in the other, this

requirement shall be considered applicable, • if a requirement is differently expressed in the two forms, it shall be considered as two

different requirements and both of them shall be considered applicable unless they are in contradiction (i.e. both cannot be simultaneously fulfilled). In this case, the requirement given in the Doors file shall be considered applicable.




2 Documents

2.1 Applicable Documents [AD 01] EMC/ESD Design and Test Requirements GAL-REQ-ASTD-SA-R-0004 [AD 02] Space Environment Requirements GAL-REQ-ASTD-SA-R-0005 [AD 03] General Design and Interface Requirements (GDIR) GAL-REQ-ASTD-SA-R-0002 [AD 04] SS PA Req. for Subcontractors and Suppliers GAL-RQS-GLI-SYST-A-0105 [AD 05] Satellite Thermal ICD GAL-ICD-ASTD-SA-R-0020

2.2 Reference Documents [RD 01] deleted [RD 02] Space Segment and Satellite Mission Analysis GAL-AN-ASTD-SS-R-0001 [RD 03] deleted [RD 04] Soyuz Launcher User manual Doc. No. ST-GTD-SUM-01 [RD 05] Proton Launcher User manual Doc. No. LKEB-9812-1990 [RD 06] Ariane Launcher User manual No document number [RD 07] Zenit-2 Launch Vehicle User´s Guide No document number [RD 08] Space Engineering Verification ECSS-E-10-02A [RD 09] Space Engineering Testing ECSS-E-10-03A [RD 10] Space Engineering Mechanical Part 1 Thermal Control ECSS-E-10-30 1A [RD 11] Space Engineering Mechanical Part 2 Mechanical ECSS-E-10-30 2A [RD 12] SS Design Development and Verification Plan GAL-PL-ASTD-SS-A-0003 [RD 13] Galileo Satellite Requirements Document GAL-REQ-ASTD-SA-R-0001

2.3 Acronyms AD Applicable Document AIV Assembly, Integration and Verification BOL Beginning Of Life CDR Critical Design Review CFRP Carbon Fibre Reinforced Plastic CLA Coupled Loads Analysis CMCU Clock Monitoring and Control Unit CSS Coarse Sun Sensor EAM Earth Acquisition Mode EAP Etage d’Acceleration a Poudre (Ariane 5 solid rocket booster) ECA Etage Cryotechnique d’Appoint (Ariane 5 cryogenic upper stage) EMC Electromagnetic Compatibility EOL End Of Life EPC Etage Principal Cryotechnique (Ariane 5 cryogenic main core stage)




EQM Engineering and Qualification Model ES Earth Sensor ESD Electrostatic Discharge FDV Fill and Drain Valve FGUU Frequency Generation and Up-conversion Unit FM Flight Model FOC Full Operation Capability FSS Fine Sun Sensor FVV Fill and Vent Valve GHC Good Health Check GSE Ground Support Equipment HPA High Power Amplifier ICD Interface Control Document ICDU Integrated Control and Data handling Unit IOV In Orbit Validation ISM Intermediate Safe Mode LF Liquid Filter LRR Laser Retro Reflector LV Launch Vehicle LV Latching Valve MGSE Mechanical Ground Support Equipment MISREC Mission Receiver MLI Multi-Layer Insulation MRB Material Review Board MTR Magnetic Torquer Rod NOM Normal Mode NSGU Navigation Signal Generator Unit OCM Orbit Control Mode OMUX Output Multiplexer OPF Output Filter OSR Optical Solar Reflector P/F Platform P/L Payload PCB Printed Circuit Board PCDU Power Conditioning and Distribution Unit PDR Preliminary Design Review PFM Proto-Flight Model PFM Proto-Flight Model PFSU Platform Security Unit PHM Passive Hydrogen Maser PLSU Payload Security Unit PT Pressure Transducer QM Qualification Model QSL Quasi-static Loads RAFS Rubidium Atomic Frequency Standard RCT Reaction Control Thruster RD Reference Document




RW Reaction Wheel S/C Spacecraft SA Solar Array SADM Solar Array Drive Mechanism SAHD Solar Array Hold Down SAM Sun Acquisition Mode SAR Search And Rescue SART Search And Rescue Transponder SM (ultimate) Safe Mode SRS Shock Response Spectrum SS Space Segment SSPA Solid State Power Amplifier STM Structural and Thermal Model STM Structural and Thermal Model SW Switch TB Thermal Balance test TBC To be Confirmed TBD To be Defined TC Test Coupler TCS Thermal Control Subsystem TMM Thermal Mathematical Model TP Test Port TTC Tracking Telemetry and Command Trel. Reliability temperatures for the unit reliability calculation TV Thermal Vacuum test




3 Definitions

3.1 Launchers For the launch of the Galileo S/C the following launchers are foreseen: Ariane 5 and Proton for multiple launch as well as Soyuz and Zenit for single / dual launch, always in their relevant configurations as shown in the following table. Therefore the environmental and test data given in this document are such, that they envelope the requirements of the candidate launchers.




Launcher Launch Configuration User Manual

Ariane Fregat 6 shared in : 1 lower stage of 4 S/C 1 upper stage of 2 S/C (TBC by Arianespace)

Ariane 5 User’s Manual Issue 4 Revision 0, November 2004, Arianespace. Ariane 5 preliminary CLA (AR5 ECA-Galileo feasibility study) AE/DI/S/A n°2047/01, 10/10/2001, Arianespace. Ariane 5 Complementary CLA (for dynamic behaviour understanding) AE/DI/S/A/BER/CLB/N02-43, 12/07/2002, Arianespace.

Proton Breeze M 6 in 2 stages with 3 S/C each

(TBC by ILS)

Proton Launch System Mission Planner’s Guide LKEB-9812-1990 Issue 1 Revision 5, December 2001, ILS. Proton MBM feasibility analysis for Galileosat LKET-0110-0313, October 2001, ILS.

Soyuz 2 using dedicated dispenser or 1 using a dedicated single launch adapter (using the same satellite/launcher interface locations)

Soyuz User’s Manual ST-GTD-SUM-01 Issue 3 Revision 0, April 2001, Starsem. Soyuz-2-1B-Fregat CLA 353Π-14A14-27414-1114, 25/09/2002, Starsem.

Zenit As above Zenit-2 Launch Vehicle User’s Guide Revision 2, December 1994, M.K. Yangel Yuzhnoye State Design Office. Zenit preliminary CLA Mai 2002, M.K. Yangel Yuzhnoye State Design Office. Zenit final CLA M.K. Yangel Yuzhnoye State Design Office.

Table 1 - Launchers




3.2 Coordinate Systems

3.2.1 S/C Coordinate System The S/C axis definition is shown in the following figure.

Figure 1 S/C Coordinate System

For the multiple launch scenario, the S/C is in vertical configuration, i.e. the x-axis up and parallel to the launcher x-axis. For the single launch scenario, the S/C is in horizontal configuration, i.e. the x-axis is horizontal and hence perpendicular to the launcher x-axis. In this case the z-axis is vertical, i.e. the z-axis up and parallel to the launcher x-axis.

3.2.2 Launcher Coordinate System The launcher coordinate system is shown in the following figure. The figure shows as an example Ariane 5. However, it is valid for all launchers. The details for the launchers shall be taken from the dedicated User Manuals, [RD 04 to 7].




Figure 2 Launcher Coordinate System

3.3 Strength

ENVREQ-25 : Strength The Structure and the units shall be of adequate strength to withstand the Design Loads (as a combination of mechanical and thermal loads) without yielding, failing or exhibiting excessive deformations that can endanger the mission objectives. The following definitions apply for the derivation of the dimensioning loads:




Figure 3 Derivation of the dimensioning loads

Limit Load (LL): maximum load to be encountered in service for a given probability. The mechanical part of the LL is derived from the envelope of the launchers and launch configurations expected for Galileo Satellites either directly or indirectly and incorporate adequate engineering assumptions. Design Load (DL): is the Limit Load multiplied by the Design Factor of Safety. The design must show positive Margin of Safety for the Design Loads. Buckling Load (BL): the load at which buckling occurs. Buckling can be either local or global. Yield Load (YL): is the load that produces stresses at which the material exhibits a permanent deformation of ε = 0.002. Ultimate Load (UL): is the load that produces the maximum tensile, compressive or shear stresses that the material can sustain. Qualification Load (QL): the loads that will be applied to the item during the qualification test campaign. In general, the Qualification Loads are above the Design Loads. Proto-Qualification Load (PQL): the loads that will be applied to the item during the proto-qualification test campaign. In general, the Proto-Qualification Loads are below the Design Loads. Acceptance Load (AL): the loads that will be applied to the item during the acceptance test campaign. Factors of Safety (D, B, Y, U) : they ensure the compliance with the structural reliability objectives (D) or ensure an acceptable risk of buckling (B), yield (Y) or failure (U) during operation

LL DL YL UL

QL LL

FoSY

FoSU

AL

KA

KQ

FoSD

Design Phase

Testing Phase

FoSB

BL LL DL YL UL

QL LL

FoSY

FoSU

AL

KA

KQ / KPQ

FoSD

Design Phase

Testing Phase

FoSB

BL




Test Factors (KQ, KPQ, KA): they provide the Qualification on QM or STM (KQ), the Proto-Qualification on PFM (KPQ) or Acceptance (KA) loads at which the item has to be tested. The following definition applies for the Margin of Safety: Margin of Safety (MoS): a measurement of the distance between the Design Loads and the Yield or Ultimate Loads. MoS have to be positive. MoS = allowable loads/DLxFoS -1 Loads can be replaced by stresses if the load-stress relationship is linear.

3.4 Safety Factors

ENVREQ-44 : Safety factors The following Factor of Safety values apply for the Galileo Structure design:

Item FoSD FoSB FoSY FoSU

Metallic parts 1.4 2.0 1.25 1.5

Composite materials 1.4 2.0 - 1.5

Metallic and non metallic items not or not fully verified during STM test

1.4 3.0 2.0 3.0

Inserts and joining elements 1.4 - - 2.0

Inserts and joining elements, not or not fully verified during STM test

1.4 - - 4.0

Lifting points 1.4 - 2.0 3.0

Table 2 - Safety factors

ENVREQ-47 : Test factors for satellite structure a nd integrated satellite The following Test Factor values apply for the Galileo Structure and Satellite test phases:




Item KQ KPQ KA

Quasi-static (flight loads) 1.4 1.25 / 1.3 (*) 1.1

Quasi-static (ground loads) 1.5 (**) 1.5 (**) 1.1

Sine

Level

Sweep rate

1.25 / 1.3 (*)

2 octave/min.

1.25 / 1.3 (*)

4 octave/min.

1.0

Acoustic (level/duration) 3 dB/120 s 3 dB/60 s 0 dB/60 s

Shock 0 dB (***) 0 dB (***) NA

Random Covered by acoustic and sine

(*) 1.3 specifically applies to the Soyuz load cases identified in the tables presenting the quasi-static and sine load cases (**) Note that the qualification requirements relative to the flight quasi-static load cases cover the ground load cases (taking into account the different safety factors) (***) The qualification test margin (3 dB) is covered at unit test level

Table 3 - Test factors for satellite structure an d integrated satellite

ENVREQ-50 : Factors of Safety for units The following Factor of Safety values apply for the Galileo units design:

Item FoSD FoSB FoSY FoSU

Metallic parts 1.4 2.0 1.25 1.5

Composite materials 1.4 2.0 - 1.5

Metallic and non metallic items not or not fully verified by test

1.4 3.0 2.0 3.0

Inserts and joining elements 1.4 - - 2.0

Inserts and joining elements, not or not fully verified by test

1.4 - - 4.0

Table 4 - Factors of Safety for units

ENVREQ-53 : Test factors for units The following Test Factor values apply for the Galileo unit test phases:




Item KQ KPQ KA

Quasi-static (flight loads) 1.25 1.25 NA

Quasi-static (ground loads) 1.5 1.5 NA

Sine

Level

Sweep rate

1.25

2 octave/min.

1.25

4 octave/min.

NA

Random (level/duration)

Level

Duration

1.4 (g rms)

180 s

1.4 (g rms)

60 s

1.0 (g rms)

60 s

Acoustic (level/duration) 3 dB/180 s 3 dB/60 s 0 dB/60 s

Shock 3 dB NA NA

Note: Section 8.1.6 defines the qualification and acceptance levels applicable to the units. The unit flight levels can be derived using this table.

Table 5 - Test factors for units




4 Environment This section presents the mechanical, thermal and climatic environment the satellite and the satellite units will be submitted to. The requirements applying on the satellite and on the units are presented in separate sections.

4.1 Mechanical Environment

4.1.1 Satellite Level Mechanical Environment

4.1.1.1 Ground Operations

4.1.1.1.1 Satellite AIT and Transport Operations This section defines natural and induced mechanical environment to which the Galileo S/C hardware and the associated ground support equipment is subjected during ground operations. The ground operation phase starts with the satellite assembly and ends before launch. It includes manufacturing, assembly, integration and verification (AIV) and storage activities for the satellite.

ENVREQ-61 : Design loads for the transportation co ntainers The Transportation Containers shall be designed to survive the loads as specified in the following. Vertical and horizontal loads shall be considered as acting simultaneously (un-attenuated input into the MGSE).

Operation Applied Load Factors (g) Remark General transportation (road, air, ship)

Vertical ± 3.0

Horizontal ± 2.0

The loads are defined at the load bed of the transport vehicle

Ground transportation

± 3.0 ± 2.0

Handling/ hoisting ± 2.0 ± 1.0 Transportation shock

10 g

Duration up to 10 ms saw tooth 10 g is the maximum level which applies in any direction

Handling Shock -

-

Shock, equivalent to a fall from 100 mm height, one edge of the container remains on ground

Table 6 - Transportation Containers Limit Loads




ENVREQ-64 : Load attenuation by the transportation container The design of the transportation container shall be such that the transportation loads will be limited to the values given in the following table. Furthermore the satellite shall be designed to be submitted the following handling and transportation loads. The loads shall be applied on the S/C in clamped condition at an infinitely rigid interface. Vertical and horizontal loads shall be considered as acting simultaneously. Operation Applied Load Factor (g) Remark Handling ±2 longitudinal

±0.1 transverse Simultaneously applied

Transport ±2 2 g is the maximum level which applies in any direction

Table 7 - Satellite Transportation Limit Loads

ENVREQ-67 : Hoisting For hoisting a design safety factor of 4 shall be applied on the handling loads. Note 1: For the satellite to MGSE interface a design safety factor of 4 shall be considered. Note 2: For the overall MGSE design, a minimum design safety factor of 4 shall also be used. However, internal subcontractor safety rules may also be considered and submitted to space segment for approval.

4.1.1.1.2 Launch site operations During the Launch site operations, the spacecraft is submitted to static and low frequency dynamic loads which are defined in this section for each applicable launcher. The following values are flight level data. The safety factors of - 3.4 shall be considered.

4.1.1.1.2.1 Ariane 5

ENVREQ-71 : Integration of the Satellite on Ariane 5 The final integration of the Ariane 5 launcher is performed vertically and the launcher is transported vertically from the assembly building to the launch pad, i.e. the satellite always remains with +Xsat vertical and oriented upwards after propellant loading. The Ground loads are covered by the launch loads.

4.1.1.1.2.2 Soyuz

ENVREQ-73 : Integration of the Satellite on Soyuz (in Kourou) The final integration of the Soyuz launcher when launched from Kourou (dual launch or single launch) is performed vertically and the launcher is transported vertically from the assembly building to the launch pad. Therefore in dual launch with Soyuz, the satellite always remains with +Xsat vertical and oriented upwards after propellant loading. In single launch with Soyuz, the satellite always remains with +Zsat vertical and oriented upwards after propellant loading. The Ground loads are covered by the launch loads.




ENVREQ-74 : Integration of the Satellite on Soyuz (in Baikonur) The final integration of the Soyuz launcher when launched from Baikonur (dual launch or single launch) is performed vertically and the launcher is transported horizontally from the assembly building to the launch pad where it is erected.

ENVREQ-75 : Single launch: In this case the following ground loads apply (on all axes simultaneously): Step 1 (launcher in vertical configuration): Along +Zsat): -1 g (static) ± 0.8 g (dynamic) Along Xsat and Ysat: ± 0.4 g (dynamic) Step 2 (during launcher transportation): Along -Xsat: 1 g (static) ± 0.55 g (dynamic) Along Ysat and Zsat: ± 0.3 g (dynamic) Step 3 (during launcher erection): TBD

ENVREQ-76 : Dual launch: In this case the following ground loads apply (on all axes simultaneously): Step 1 (launcher in vertical configuration): Along +Xsat direction: -1 g (static) ± 0.8 g (dynamic) Along Ysat and Zsat: ± 0.4 g (dynamic) Step 2 (during launcher transportation): The vertical axis (V) during transportation can be any direction perpendicular to the satellite Xsat axis. Along the V direction: 1 g (static) ± 0.55 g (dynamic) Along any direction perpendicular to (V): ± 0.3 g (dynamic) Step 3 (during launcher erection): TBD

4.1.1.1.2.3 Zenit Integration of the Satellite on Zenit The final integration of the Zenit launcher (dual launch or single launch) is performed horizontally and the launcher is transported horizontally from the assembly building to the launch pad where it is erected.

ENVREQ-79 : Single Launch: The following ground loads apply (on all axes simultaneously): Step 1 (during launcher transportation): . Along Xsat: - 1g (static) +/- 0.2 g (dynamic) . Along Zsat. +/- 0.35 g (dynamic) . Along Ysat: +/- 0.2 g (dynamic) Step 2 (during launcher erection): . Along any direction between -Xsat and -Zsat: 1.35 g Step 3 (in launch configuration):




. Along any direction perpendicular to the Zsat axis: . 0.6m amplitude . Frequency within [0.25 Hz - 0.3 Hz]

ENVREQ-80 : Dual launch: The following ground loads apply (on all axes simultaneously): Step 1 (during launcher transportation): The vertical axis (V) during transportation can be any direction perpendicular to the satellite Xsat axis. . Along the V direction: - 1g (static) +/- 0.2 g (dynamic) . Along Xsat: +/- 0.35 g (dynamic) . Along the normal to V and Xsat: +/- 0.2 g (dynamic) Step 2 (during launcher erection): . Along any axis W between the V direction and the -Xsat direction: 1.35 g Step 3 (in launch configuration): . Along any direction perpendicular to the Xsat axis: . 0.6m amplitude . Frequency within [0.25 Hz - 0.3 Hz]

4.1.1.1.2.4 Proton

ENVREQ-81 : Integration of the Satellite on Proton The final integration of the Proton launcher is performed horizontally and the launcher is transported horizontally from the assembly building to the launch pad where it is erected.

ENVREQ-82 : The following ground loads apply (on all axes simultaneously): Step 1 (during launcher transportation): The vertical axis (V) during transportation can be any direction perpendicular to the satellite Xsat axis. . Along the V direction: 1 g (static) +/- 0.5 g (dynamic) . Along Xsat: +/- 0.5 g (dynamic) . Along the normal to V and Xsat: +/- 0.4 g (dynamic) Step 2 (during launcher erection): . Along any axis W between the V direction and the -Xsat direction: 1 g (static) +/- 0.5 g (dynamic) . Along any axis perpendicular to the W axis: +/- 0.15 g (dynamic)

4.1.1.2 Launch Mechanical Environment During launch and ascent, the spacecraft is subjected to static and dynamic loads induced by the launch vehicle via the dispenser. Such excitation may be due either to aerodynamic (wind, gusts, buffeting at transonic velocity), to the propulsion system (longitudinal acceleration, thrust build-up or tail-off transients, structure-propulsion coupling, etc.) or to the launcher attitude control system. Furthermore, the spacecraft is submitted to acoustic noise resulting from the launcher propulsion and to shocks generated by several launch events such as fairing separation and satellite separation.




The quasi-static and sine loads given in this document consider the horizontal (single launch) as well as the vertical launch configuration of the S/C. For the horizontal case the loads are taken directly from [RD 04] and [RD 07] for Soyuz and Zenit. General: The dispenser (including the separation system) is in the responsibility of the launcher authority. The external loads which are applicable for the S/C are given at the interface dispenser to S/C and are provided by the launcher authority (the values herein are derived from preliminary CLAs, performed by the launcher authorities).

4.1.1.2.1 Quasi- Static Loads During flight, low frequency dynamic and steady loads are combined to produce quasi-static loads (QSL).

ENVREQ-84 : Quasi Static Design Load The QSL which shall be considered for the dimensioning of the S/C are given in the following table, which contains the dimensioning accelerations of the applicable launchers. All loads in the figure below are TBC, since the values are derived partially from preliminary CLAs, which were performed with a S/C model which does not fully represent the present S/C design and which were performed with draft dispenser design.

Longitudinal Lateral Sized by -2,7 2 -2,7 -2

AR5 maximum dynamic pressure

-4,7 1 -4,7 -1

AR5 SRB end of flight

-3 1,2 -3 -1,2 3 1,2 3 -1,2

Proton stage 1/2 separation

-5,3 (S) -0,6 (S) -5,3 (S) 0,6 (S)

Soyuz max nxi

Table 8 - Quasi-static Flight Limit Loads (envelo pe for multiple launch configuration)




Longitudinal Lateral Sized by 1 1 1 -1

Zenith stages separation

-1,3 (S) 0,8 (S) -1,3 (S) -0,8 (S)

Soyuz 2nd/3rd stage separation

-1,6 (S) 1,8 (S) -1,6 (S) -1,8 (S)

Soyuz lift-off

-1,95 1,5 -1,95 -1,5

Zenith lift-off

-4,5 1 -4,5 -1

Zenith LV flight

-5 (S) 0,5 (S) -5 (S) -0,5 (S)

Soyuz 1st stage max acc

Table 9 - Quasi-static Flight Limit Loads (envelo pe for single launch configuration)

Notes: • These values are given in the longitudinal and lateral axes of the launch vehicle.

Transformation in the spacecraft axes shall be performed as indicated in section 3.2.1 according to the launch configuration (multiple or single launch).

• The positive values represents tension, the negative values represents compression. • Lateral loads may act in any direction simultaneously with longitudinal loads. • Thermo-elastic cases shall not be applied simultaneously except for cases to be

agreed with the space segment prime. • The QSL values and angular acceleration values apply at the CoG of the S/C when it

is rigidly clamped at its four interfaces (gravity load is included) • The values with mention (S) correspond to Soyuz load cases for which dedicated test

factors apply as specified in section 3.4.

4.1.1.2.2 Sine Vibrations

ENVREQ-97 : Sine Vibrations The sine vibration flight levels at the dispenser / spacecraft interface are presented in the table below for the multiple launch configuration and for the single launch configuration. The spectrum takes into account all sinusoidal or transient vibrations in this bandwidth and covers all applicable launchers. For multiple launch configuration:

Frequency range 4-5 Hz 5-20 Hz 20-30 Hz 30-100 Hz Longitudinal Sine Level 13,9 mm 1.4g 1g (S) 1g

Table 10 - Sine longitudinal envelope (flight lev el)




Frequency range 2-5 Hz 5-30 Hz 30-60 Hz 60-100 Hz Lateral Sine Level 8 mm 0,8g (S) 0,6g (S) 0,6g

Table 11 - Sine lateral envelope (flight level)

For single launch configuration: Frequency range 4-5 Hz 5-10 Hz 10-30 Hz 30-60 Hz 60-100 Hz

Longitudinal Sine Level 5 mm 0.5g (S) 1g (S) 0.6g (S) 0.3g (S)

Table 12 - Sine longitudinal envelope (flight lev el)

Frequency range 2-5 Hz 5-30 Hz 30-60 Hz 60-100 Hz Lateral Sine Level 8 mm 0,8g (S) 0,6g (S) 0,2g (S)

Table 13 - Sine lateral envelope (flight level)

Notes: - These values are given in the longitudinal and lateral axes of the launch vehicle. Transformation in the spacecraft axes shall be performed as indicated in section 3.2.1 according to the launch configuration (multiple or single launch). - The values with mention (S) correspond to Soyuz load cases for which dedicated test factors apply as specified in section 3.4.

ENVREQ-111 : Test factors for sine- testing For qualification, proto-qualification and acceptance sine environment testing, the test factors and the sweep rates as specified in section 3.4 shall be applied.

4.1.1.2.3 Random Vibration Random vibrations at the S/C are generated by propulsion system operation and by the adjacent structure vibro-acoustic response. Maximum excitation levels are obtained during the flight of the first stage. It is expected that the random environment be covered by the acoustic (high frequencies) and by the sine (low frequencies) environments as specified in the next sections. This shall be demonstrated through the STM test sequence.

4.1.1.2.4 Acoustic Noise The following acoustic environment is the envelope of the acoustic environment of all applicable launchers.

ENVREQ-114 : Flight level acoustic noise spectrum The following Flight Level acoustic noise spectrum applies:




Octave Band Frequency Hz

Acoustic Level dB

31.5 128 63 132 125 136 250 136 500 134 1000 129 2000 125 Overall Level 141.4

Table 14 - Flight Level acoustic noise spectrum

ENVREQ-117 : Acoustic test factor and durations For qualification, proto-qualification and acceptance acoustic environment testing, the test factors and the test durations as specified in section 3.4 shall be applied.

4.1.1.2.5 Shock The spacecraft is subjected to external shocks mainly during separation of the fairing and on actual separation of the spacecraft. Internal shock sources coming from the solar array deployment can be neglected as for the solar array deployment low shock systems (thermal knifes) will be used.

4.1.1.2.5.1 Launcher separation shock

ENVREQ-120 : Launcher separation induced shock The separation shock flight level which shall be considered at the S/C to separation system interface is defined in the table and figure below. It covers the envelope of the spacecraft separation environment (which encloses shock spectra of the applicable launch systems Ariane 5, Proton, Soyuz and Zenit).

Frequency (Hz) Acceleration (g) 100 100 600 1800 2000 5000

10 000 5000

Table 15 - Launcher separation induced shock




Shock spectrum

100

1000

10000

100 1000 10000

Acceleration (g)

Figure 4 Launcher separation shock spectrum (flight level)

4.1.1.2.5.2 Solar Array Deployment Shock

ENVREQ-127 : Solar array release shock For the solar array release and deployment system, low shock systems (thermal knifes) will be used. Therefore the induced shock levels are negligable.

4.1.2 Unit Level Mechanical Environment The mechanical environment defined in the present section applies to all satellite units. It has been defined taking into account the environment the satellite is submitted to and the selected satellite accommodation. It covers the transportation, handling, storage, pre-launch, launch/ascent, transfer orbit and in-orbit environments. Where specific levels are not stated in this section, then the environmental design requirements are those specified in the test section for unit's qualifications levels.

ENVREQ-131 : General design rule for units Units shall be designed to achieve their specified performance requirements during and/or after exposure to the specified environments. Generally, the demonstration can be done by test or analysis after space segment approval unless demonstration by test is clearly specified as stated in section 8.1.

ENVREQ-132 : If handling, transportation and storage environments drive the design of a unit, then the space segment prime shall be specifically informed.




4.1.2.1 Handling, Transportation and Storage

ENVREQ-134 : On ground modes The normal modes of transportation to be considered shall include air, sea and overland. The environments experienced by the unit during the fabrication, delivery, storage and installation phases shall be controlled so as to be significantly less severe than the ones experienced during launch and ascent conditions.

4.1.2.1.1 Handling

ENVREQ-136 : Handling loads The units shall be designed to withstand the ground handling limit loads as defined in the following table. In each case, vertical and horizontal loads can apply simultaneously. Values for the vertical condition include normal gravity (1 "g"). These loads have to be considered for any unit orientation. Operation Applied Load Factor (g) Remark Hoisting at spacecraft level ±2 vertical

±0.1 horizontal Simultaneously applied

Handling ±1.5 In all directions

Table 16 - Handling loads

4.1.2.1.2 Transport

ENVREQ-140 : Transportation loads The units shall be designed to withstand the transport limit loads as defined in the following table. In each case, vertical and horizontal loads can apply simultaneously. Values for the vertical condition include normal gravity (1 "g"). These loads have to be considered for any unit orientation. Operation Applied Load Factor (g) Remark Air transportation ±2.5 vertical

±3.5 horizontal Simultaneously applied

Ground transportation ±3.0 vertical ±2.0 horizontal

Simultaneously applied

Table 17 - Transportation loads

ENVREQ-143 : Attenuation of ground loads by the tr ansportation container Unit transportation containers and their method of transportation shall be such as to ensure that levels experienced by the units are limited to constant acceleration less severe than those specified in the previous table.




ENVREQ-144 : Attenuation of ground shock by the tr ansportation container Unit transportation containers shall be capable to be submitted to the shock defined as a drop of 100 mm onto concrete of one corner of the container, another corner lying on the concrete floor. Furthermore, the transportation containers shall be designed to ensure that the flight unit contained within is protected from and shall be undamaged by shocks.

4.1.2.1.3 Storage

ENVREQ-146 : Global storage requirement By appropriate environmental control and maintenance to be defined by the unit supplier, it shall be possible to store all elements of the spacecraft for the storage period duration of 5 years. Separate storing of critical items requiring periodic checking and/or storage in a particular orientation, storing of piece parts rather than integrated units or assemblies, and the replacement of items demonstrated to be subject to excessive deterioration, will be allowed during the storage period. These requirements have to be clearly defined by the unit supplier in the unit user's manual.

4.1.2.1.4 Vacuum

ENVREQ-148 : Vacuum testing Satellite units will be submitted to vacuum during the satellite thermal vacuum test and shall be able to be operated at pressure lower or equal to 10-3 hPa.

4.1.2.2 Pre-launch and launch/ascent environment

4.1.2.2.1 General The environments of pre-launch and launch/ascent phases are covered by test specifications of section 8.1. These tests cover quasi-static, transient, sine, shock, random and acoustic environment of the launch phase.

4.1.2.2.2 Shocks Shocks will be experienced at fairing separation, at satellite separation and when the solar array deployments devices are activated. Resultant structural response characteristics resemble the form of complex decaying sinusoid which can excite resonant responses in the unit.

ENVREQ-152 : The units shall therefore withstand the Shock Response Spectrum (SRS) of section 8.1, which depends on their location on the satellite.




4.1.2.2.3 Depressurization

ENVREQ-154 : Depressurization requirements For the design of the units, the following assumptions have to be used: • maximum pressure decay rate of 70 mbars / sec. • pressure decay of 0.8 bar in 27 sec A demonstration by test (see section 8.1) is required.

4.1.2.2.4 Corona

ENVREQ-156 : Units ON during launch will be submitted to all pressure values lower than 1 bar. They have to be designed to operate in such an environment without any arcing (see section 8.1).

4.1.2.2.5 Aero-thermal flux

ENVREQ-158 : Externally mounted units, shall be able to withstand a nominal aero-thermal flux of 1135 W/m². In addition, for some particular launch configuration, they shall be able to withstand 1500 W/m² during 70 sec.

4.1.2.2.6 Combined Thermal and Launch Loads

ENVREQ-160 : It shall be investigated, whether during launch a combination of launch loads and thermal deflections may lead to a dimensioning load case for the satellite or external units, e.g. the solar arrays, which then needs to be analysed accordingly.

4.1.2.3 In Orbit Phase

4.1.2.3.1 Quasi- Static Loads

ENVREQ-163 : Accelerations due to orbital manoeuvr es During orbit attitude correction manoeuvres, units shall be able to withstand the following accelerations: linear acceleration: 0.05 m/s² angular acceleration: 0.02 rd/s²

ENVREQ-164 : These accelerations can act simultaneously and shall be applied at the CoG of the unit.




ENVREQ-165 : Furthermore, all units (except the clocks) shall be able to meet their performance requirements under the above environment.

ENVREQ-166 : Accelerations due to orbital manoeuvr es valid for SADM In addition, during orbital manoeuvres, loads (TBD) will be induced e.g. by the solar arrays to the SADM which shall be considered in the design.

ENVREQ-167 : Angular rates of accelerations When in orbit, the satellite units shall be able to be submitted to the following angular rates: • 10°/s on any axis after separation from the launch er • 3°/s on any axis afterwards •

4.1.2.3.2 Micro-vibrations

ENVREQ-169 : Micro-vibrations are generated by the satellite actuators, mainly the reaction wheels. All units shall be able to operate and be compliant with their performance requirements under the micro-vibration environments defined in AD 03.

4.1.2.3.3 Vacuum

ENVREQ-171 : Satellite units will be submitted to the hard vacuum when in orbit. However, the units shall be able to operate nominally at any pressure below 0,15 hPa.

4.2 Thermal Environment

4.2.1 General Requirements

4.2.1.1 Operating temperature limits

ENVREQ-175 : The operating temperature limits specify the temperature extremes at and between which the unit shall be designed to operate whilst meeting all the functional requirements.

ENVREQ-176 : Operating temperature limits are applicable when an internal- or external unit is defined 'on' for the appropriate orbit mode.




ENVREQ-177 : The thermal control subsystem (TCS) shall guarantee that minimum and maximum operating temperatures during all mission phases are within the specified design limits of each unit.

ENVREQ-178 : For units having temperature stability requirements, the units shall be designed to provide the relevant performance requirements within the specified temperature variation range.

ENVREQ-179 : For units having temperature stability requirements, the thermal control subsystem (TCS) shall guarantee that the maximum temperature variation over the specified time span will never be exceeded during all the nominal mission phase.

4.2.1.2 Non- operating temperature limits

ENVREQ-181 : The Non-Operating temperature limits specify the temperature extremes at and between which the unit shall be designed to survive without functioning for any period of the mission and without performances degradation.

ENVREQ-182 : Non-Operating temperature limits are applicable when an internal- or external unit is defined 'off' for the appropriate orbit mode.

ENVREQ-183 : The thermal control subsystem shall guarantee that minimum and maximum non-operating temperatures during all mission phases, where non-operating conditions are applicable, are within the specified design limits of each unit.

4.2.1.3 Start-up temperature

ENVREQ-185 : The minimum start-up temperature specifies the minimum temperature for the unit where power can be applied, or be activated, and a unit can be switched from the non-operating mode to the operating mode, and functions nominally, when the unit temperature is brought back to the relevant operating mode temperature. The unit can be switched-on at the start-up temperature and above this temperature without suffering any damage. Between start-up temperature and minimum operating temperature the unit will operate correctly, however, full performance is not required.




4.2.1.4 Ground storage temperature

ENVREQ-187 : The ground storage temperature limits specify the temperature extremes for which the unit shall be designed to survive in a not-powered state during manufacturing, integration, handling, storage and transport.

4.2.1.5 Temperature margin and test philosophy for TCS and Units

ENVREQ-189 : The following figure specifies the definition of terms applicable for thermal design and unit thermal tests.

Figure 5 Definition of terms applicable for thermal design

ENVREQ-192 : Qualification margin (QM) The qualification margin is the temperature difference between qualification temperature limit and acceptance temperature limit. The TCS qualification margin shall be at least 5°C. It shall be extended at both temperature extremes.




ENVREQ-193 : Acceptance Margin (AM) The acceptance margin is the temperature difference between acceptance temperature limit and design temperature limit. The TCS acceptance margin shall be at least 5°C. It shall be extended at both temperature extremes.

ENVREQ-194 : Design Margin (DM) The design margin is the temperature difference between design temperature limit and predicted temperature limit. The TCS design margin shall be positive at both temperature extremes.

ENVREQ-195 : Uncertainty Margin (UM) UM is a lack of certitudes resulting from inaccuracies of input parameters (material properties, environmental data) and or analysis process (modelling assumptions). The uncertainty margin is the temperature difference between design temperature limit and calculated temperature limit. The TCS uncertainty margin shall be at least 10°C f or internal elements. I shall be extended at both temperature extremes. The TCS uncertainty margin shall be at least 15°C f or external elements. I shall be extended at both temperature extremes. The uncertainty margin shall be evolving with project maturity (PDR, STM-TB-test, CDR, PFM-TB-test) according to ECSS-E-30 Part 1A and on justified sensitivity analyses. The uncertainty margin will be defined by TCS and agreed by Prime. Note on clock and battery temperature range: For the clocks (PHM and RAFS), the narrow design temperature range does not allow a 10 °C uncertainty margin in hot and cold case to be co nsidered. The clocks will be nominally regulated at 5 °C for the normal orbit case. An uncertainty margin of 2°C shall be considered. Furthermore the thermal design shall provide a minimum margin of 15% between maximum needed heater power and effectively installed heater power. For the battery, the narrow design temperature range does not allow either a 10 °C uncertainty margin in hot and cold case to be considered. Specifically for the battery, the temperature margin between the predicted temperature and the design temperature shall then be equal to 5 °C in hot case and 3 °C in cold case.

ENVREQ-196 : Unit Acceptance Temperature Range: The unit supplier shall perform acceptance tests with his unit under acceptance operating temperature limits, with functioning, fulfilling all required performances and without performance degradation.

ENVREQ-197 : The unit supplier shall perform acceptance tests with his unit under acceptance non-operating temperature limits, without functioning and without performance degradation.




ENVREQ-198 : The unit supplier shall perform qualification tests with his unit under qualification operating temperature limits, with functioning, fulfilling all required performances and without performance degradation.

ENVREQ-199 : The unit supplier shall perform qualification tests with his unit under qualification non-operating temperature limits, without functioning and without performance degradation.

ENVREQ-200 : The unit supplier shall perform during his acceptance or qualification tests with his unit the application of the start-up temperature without performance degradation.

ENVREQ-201 : The unit supplier shall perform qualification tests with his unit under ground storage temperature, without functioning and without performance degradation.

4.2.2 Mission Overview

4.2.2.1 Lifetime

ENVREQ-204 : The lifetime of each Satellite (whether part of the operational constellation or as in-orbit spare) shall be a minimum of 12 years. This lifetime excludes any periods of on-ground storage and also excludes the AIT phase up to completion of spacecraft acceptance testing on ground.

4.2.2.2 Satellite Phases with S/C Modes The satellite follows different 'satellite phases'. The satellite follows a defined sequence of operations and achieves defined conditions, which are reflected by corresponding satellite modes i.e. during a satellite phase the satellite can be in one satellite mode or can pass through a sequence of satellite modes. The Satellite Phases and their correlation to Satellite modes are defined below. Note: Safe Hold Phase includes ultimate and intermediate safe modes.




Figure 6 Satellite Phases




ENVREQ-208 : The satellite thermal design shall provide thermally acceptable environments (temperatures, temperature gradients and temperature stability) for all its units and components during its entire lifetime and all the satellite modes. The correlation of the main satellite phases and its applicable S/C modes is shown in the table below: Satellite Phase S/C Mode

Ground Storage & Transport Phase

Pre-Launch Phase Launch Phase Deployment Phase In Orbit Test Phase

Mission Operating Phase Spare Phase

Orbit Keeping Phase

Repositioning Manoeuvre Phase

Safe Hold Phase

Graveyard Phase

Off Mode=Standby Mode X X

Test Mode X X

Launch pad to fairing jettison X

Fairing jettison to separation X

Init Mode X

Sun Acquisition Mode X

Earth Acquisition Mode X X X

Normal Mode X X X X X X

Orbit Change Mode X X X X X

Safe Mode X

Intermediate Safe Mode X

Table 18 - Correlation of the main satellite phas es and its applicable S/C modes

ENVREQ-211 : The satellite modes and the mode transitions which shall be analysed and taken into account by the S/C Thermal Control Subsystem are shown in the figure below:




OFFMode

STANDBYMode

LAUNCHMode

INITSequence

SUNACQUISITION

Mode

EARTHACQUISITION

Mode

NORMALMode

SAFEMode

ORBITCHANGE

Mode

IntermediateSAFEMode

Umbilical orSatellite Power

(int. or ext.power)

S/C Separation

TC

autom.

autom.

TC

TC

TC

TC

TC TC

TCTC

TC

TC

Severe FailureFDIR Level 2+3

Sev

ere

Fai

lure

FD

IR L

eve

l 4

Seve

re F

ailu

re

FDIR

Lev

e l 2

+3+

4

Severe Failure

FDIR Level

2+3+4

Se v

e re

Fa i

lure

FD

IR L

eve

l2+

3+4

Severe FailureFDIR Level

2+3+4

After ICDUReconfigurationFDIR Level 2 + 3

After ICDUReconfigurationFDIR Level 2+3

Failure during INIT

Nominal transition

Exceptional transition

Alarm/Failuretransition

TC

TC (tbc)After ICDU

Reconfiguration

FDIR Level 3

Umbilical unpluggedSatellite Power

= int.power

FDIR Level 4inhibited by Gnd TC

TC

TCTC

TC

Figure 7 Satellite modes




4.2.2.3 Platform unit's configuration

ENVREQ-215 : The platform unit's configuration cases to be considered by the thermal control are presented in the following tables for the Ground Phase (Off Mode and Stand By/Test Mode), for the Launch Phase (Launch Pad to Fairing Jettison, Fairing Jettison to Separation, INIT sequence) and for the In Orbit Phase (SAM, EAM, OCM, NOM, ISM and SM Modes):

Qty Qty Temperature Qty Temperature Qty Temperature Qty Temperature Qty TemperatureOn To guarantee On To guarantee On To guarantee On To guarantee On To guarantee

TX/RX1 S-Band Transponder (M) 1 0 storage 1 T-op 1 T-op 1 T-op 1 T-opTX/RX2 S-Band Transponder R ( Note 3)) 1 0 storage 1 T-op 1 T-op 1 T-op 1 T-opTTCHYB Hybrid 1 *) storage *) T-op *) T-op *) T-op *) T-opTTCANTN S-Band Antenna Nadir 1 *) storage *) T-op *) T-op *) T-op *) T-opTTCANTZ S-Band Antenna Zenith 1 *) storage *) T-op *) T-op *) T-op *) T-opTTCHAR RF cables 1 *) storage *) T-op *) T-op *) T-op *) T-op

LRR LRR Laser retro reflector 1 *) storage *) T-op *) T-op *) T-op *) T-op

PFSU PFSU Platform Security Unit (hot red.) 1 0 storage 1 T-op 1 T-op 1 T-op 1 T-op

PCDU Power Conditioning and Distribution Unit 1 0 storage 1 T-op 1 T-op 1 T-op 1 T-opSADM-Y Solar Array Drive Mechanism -Y 1 0 storage 1 T-op 0 Tnon-op 0 Tnon-op 0 Tnon-opSADM+Y Solar Array Drive Mechanism +Y 1 0 storage 1 T-op 0 Tnon-op 0 Tnon-op 0 Tnon-opBATT Battery 1 0 storage 1 T-op 1 T-op 1 T-op 1 T-op

ICDU Integrated Control and Data Handling Unit 1 0 storage 1 T-op 1 T-op 1 T-op 1 T-opFSS1 Fine Sun Sensor (-x) 1 0 storage 1 T-op 0 Tnon-op 0 Tnon-op 1 T-opFSS2 FineSun Sensor (-x) 1 0 storage 1 T-op 0 Tnon-op 0 Tnon-op 0 Tnon-opCSS1 Coarse Sun Sensor (+z) 1 0 storage 1 T-op 0 Tnon-op 0 Tnon-op 1 T-opCSS2 Coarse Sun Sensor (-z) 1 0 storage 1 T-op 0 Tnon-op 0 Tnon-op 1 T-opES1 Infrared Earth Sensor 1 1 0 storage 1 T-op 0 Tnon-op 0 Tnon-op 0 Tnon-opES2 Infrared Earth Sensor 2 1 0 storage 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-opGYRO GYRO 1 0 storage 1 T-op 0 Tnon-op 0 Tnon-op 1 T-opMTR1 Magnetic Torquer 1 1 0 storage 1 T-op 0 Tnon-op 0 Tnon-op 0 Tnon-opMTR2 Magnetic Torquer 2 1 0 storage 1 T-op 0 Tnon-op 0 Tnon-op 0 Tnon-opRW L1 -4 Reaction W heels 1 - 4 4 0 storage 3/4**) T-op 0 Tnon-op 0 Tnon-op 3/4**) T-op

PROTank Tank 1 *) storage *) T-op *) T-op *) T-op *) T-opPROFVV Fill and Vent Valve 1 *) storage *) T-op *) T-op *) T-op *) T-opPROPT Pressure Transducer 1 0 storage 1 T-op 0 Tnon-op 0 Tnon-op 1 T-opPROLF Filter 1 *) storage *) T-op *) T-op *) T-op *) T-opPROFDV Fill and Drain Valve 1 *) storage *) T-op *) T-op *) T-op *) T-opPROLVA Latching Valve 1 1 *) storage *) T-op *) T-op *) T-op *) T-opPROLVB Latching Valve 2 1 *) storage *) T-op *) T-op *) T-op *) T-opPROTPA Test Port A 1 *) storage *) T-op *) T-op *) T-op *) T-opPROTPB Test Port B 1 *) storage *) T-op *) T-op *) T-op *) T-opPRORCT1A Reaction Control Thruster 1A 1 0 storage 0 Tnon-op 0 Tnon-op 0 Tnon-op 1 T-opPRORCT2A Reaction Control Thruster 2A 1 0 storage 0 Tnon-op 0 Tnon-op 0 Tnon-op 1 T-opPRORCT3A Reaction Control Thruster 3A 1 0 storage 0 Tnon-op 0 Tnon-op 0 Tnon-op 1 T-opPRORCT4A Reaction Control Thruster 4A 1 0 storage 0 Tnon-op 0 Tnon-op 0 Tnon-op 1 T-opPRORCT1B Reaction Control Thruster 1B 1 0 storage 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-opPRORCT2B Reaction Control Thruster 2B 1 0 storage 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-opPRORCT3B Reaction Control Thruster 3B 1 0 storage 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-opPRORCT4B Reaction Control Thruster 4B 1 0 storage 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op

PFHAR Platform Power Harness 1 *) storage *) T-op *) T-op *) T-op *) T-opPLHAR Payload Power Harness 1 *) storage *) T-op *) T-op *) T-op *) T-op

Fairing Jettison to Separation

INIT Sequence on ground

S/C ModeOff Mode

storage&transport

AOCS

EPS

Stand By/Test Mode Launch Pad to Fairing Jettison

TTC

Item Equipment

PRPS

Harness

T-op = operating TemperatureTnon-op = non-operating Temperature

Note2: Battery requirements vary for launch and eclipse conditionsNote 3): TTC transponder redundant (Rx=on; Tx=off).

3/4**): Three wheels for miniumum , four for maximum dissipation

Note1: Transitions form modes may imply start-up temperature requirements

*): passive element




Qty Qty Temperature Qty Temperature Qty Temperature Qty Temperature Qty Temperature Qty TemperatureOn To guarantee On To guarantee On To guarantee On To guarantee On To guarantee On To guarantee

TX/RX1 S-Band Transponder (M) 1 1 T-op 1 T-op 1 T-op 1 T-op 1 T-op 1 T-opTX/RX2 S-Band Transponder R ( Note 3)) 1 1 T-op 1 T-op 1 T-op 1 T-op 1 T-op 1 T-opTTCHYB Hybrid 1 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-opTTCANTN S-Band Antenna Nadir 1 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-opTTCANTZ S-Band Antenna Zenith 1 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-opTTCHAR RF cables 1 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-op

LRR LRR Laser retro reflector 1 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-op

PFSU PFSU Platform Security Unit (hot red.) 1 1 T-op 1 T-op 1 T-op 1 T-op 1 T-op 1 T-op

PCDU Power Conditioning and Distribution Unit 1 1 T-op 1 T-op 1 T-op 1 T-op 1 T-op 1 T-opSADM-Y Solar Array Drive Mechanism -Y 1 1 T-op 1 T-op 1 T-op 1 T-op 1 T-op 1 T-opSADM+Y Solar Array Drive Mechanism +Y 1 1 T-op 1 T-op 1 T-op 1 T-op 1 T-op 1 T-opBATT Battery 1 1 T-op 1 T-op 1 T-op 1 T-op 1 T-op 1 T-op

ICDU Integrated Control and Data Handling Unit 1 1 T-op 1 T-op 1 T-op 1 T-op 1 T-op 1 T-opFSS1 Fine Sun Sensor (-x) 1 1 T-op 1 T-op 1 T-op 1 T-op 1 T-op 1 T-opFSS2 FineSun Sensor (-x) 1 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-opCSS1 Coarse Sun Sensor (+z) 1 1 T-op 1 T-op 1 T-op 1 T-op 1 T-op 1 T-opCSS2 Coarse Sun Sensor (-z) 1 1 T-op 1 T-op 1 T-op 1 T-op 1 T-op 1 T-opES1 Infrared Earth Sensor 1 1 0 Tnon-op 1 T-op 1 T-op 1 T-op 0 Tnon-op 1 T-opES2 Infrared Earth Sensor 2 1 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-opGYRO GYRO 1 1 T-op 1 T-op 0 Tnon-op 1 T-op 1 T-op 0 Tnon-opMTR1 Magnetic Torquer 1 1 0 Tnon-op 0 Tnon-op 1 T-op 0 Tnon-op 0 Tnon-op 1 T-opMTR2 Magnetic Torquer 2 1 0 Tnon-op 0 Tnon-op 1 T-op 0 Tnon-op 0 Tnon-op 1 T-opRWL1 -4 Reaction Wheels 1 - 4 4 3/4**) T-op 3/4**) T-op 3/4**) T-op 3/4**) T-op 3/4**) T-op 3/4**) T-op

PROTank Tank 1 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-opPROFVV Fill and Vent Valve 1 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-opPROPT Pressure Transducer 1 1 T-op 1 T-op 1 T-op 1 T-op 1 T-op 1 T-opPROLF Filter 1 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-opPROFDV Fill and Drain Valve 1 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-opPROLVA Latching Valve 1 1 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-opPROLVB Latching Valve 2 1 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-opPROTPA Test Port A 1 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-opPROTPB Test Port B 1 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-opPRORCT1A Reaction Control Thruster 1A 1 1 T-op 1 T-op 0 Tnon-op 1 T-op 1 T-op 0 Tnon-opPRORCT2A Reaction Control Thruster 2A 1 1 T-op 1 T-op 0 Tnon-op 1 T-op 1 T-op 0 Tnon-opPRORCT3A Reaction Control Thruster 3A 1 1 T-op 1 T-op 0 Tnon-op 1 T-op 1 T-op 0 Tnon-opPRORCT4A Reaction Control Thruster 4A 1 1 T-op 1 T-op 0 Tnon-op 1 T-op 1 T-op 0 Tnon-opPRORCT1B Reaction Control Thruster 1B 1 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-opPRORCT2B Reaction Control Thruster 2B 1 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-opPRORCT3B Reaction Control Thruster 3B 1 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-opPRORCT4B Reaction Control Thruster 4B 1 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op

PFHAR Platform Power Harness 1 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-opPLHAR Payload Power Harness 1 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-op

Note2: Battery requirements vary for launch and eclipse conditionsNote 3): TTC transponder redundant (Rx=on; Tx=off).

3/4**): Three wheels for miniumum , four for maximum dissipation *): passive element

Note1: Transitions form modes may imply start-up temperature requirements

PRPS

Harness

T-op = operating TemperatureTnon-op = non-operating Temperature

AOCS

EPS

Sun Acquisition ModeEarth Acquisition Mode

TTC

S/C Mode

Item Equipment

Safe Mode Intermediate Safe ModeNormal Mode Orbit Change Mode

Table 19 - Platform unit's configuration

4.2.2.4 Payload unit's configuration

ENVREQ-220 : The payload unit's configuration cases to be considered by the thermal control are presented in the following tables for the Ground Phase (Off Mode and Stand By/Test Mode), for the Launch Phase (Launch Pad to Fairing Jettison, Fairing Jettison to Separation, INIT sequence) and for the In Orbit Phase (SAM, EAM, OCM, NOM, ISM and SM Modes):




Qty Temperature Qty Temperature Qty Temperature Qty Temperature Qty Temperature

on To guarantee on To guarantee on To guarantee on To guarantee on To guarantee

RAFS1 Rubidium Atomic Frequency Standard 1 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

RAFS2 Rubidium Atomic Frequency Standard 2 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

Passive Hydrogen M aser 1 (***) 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

Ion pump (part of PHM 1) 1 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

Passive Hydrogen M aser 2 (***) 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

Ion pump (part of PHM 2) 1 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

CM CU Clock M onitoring and Control Unit 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

PLSU Payload Securit y Unit 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

NSGU Navigation Signal Generator Unit 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

FGUU Frequency Generation and Upconversion Unit 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

IPTC1-3 Input Test Coupler 1-3 3 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

SPLIT1-3 Low Power HB Split ter 1-3 3 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

NAVHPA1 SSPA HB1 (L1)2 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

NAVHPA2 SSPA HB2 (L1)2 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

NAVHPA3 SSPA HB1/HB2 Spare (L1)2 1 0 storage 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op

NAVHPA4 SSPA LB1 (E5)2 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

NAVHPA5 SSPA LB1 Spare (E5)2 1 0 storage 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op

NAVHPA6 SSPA LB2 (E6)2 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

NAVHPA7 SSPA LB2 Spare (E6)2 1 0 storage 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op

NAVSW1-3 RF High Power Switch 1-3 3 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

NAVLOAD1-3 RF High Power Load 3 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

OM UX OM UX Diplexer LB (E5/E6) 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

OPF 1 Output Filter 1 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

OPF 2 Output Filter 2 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

OPTC1-3 Output Test Couplers 1-3 3 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

NAVANT L-Band Navigat ion Antenna 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

M ISREC C-Band Receiver 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

M ISANT C-band Antenna Array 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

SARANT SAR Rx/Tx Antenna 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

SARIPTC SAR Test Coupler 406 M Hz 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

SAROPTC SAR Test Coupler L-Band 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

SART SAR Transponder Assembly 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

RTU Remote Terminal Unit 1 0 storage 1 T-op (*) 0 Tnon-op 0 Tnon-op 0 Tnon-op

HPHAR01-17 RF Cable 1-17 17 *) storage *) T-op *) T-op *) T-op *) T-op

LPHAR RF LP Harness - *) storage *) T-op *) T-op *) T-op *) T-op

CBHAR C-Band Harness - *) storage *) T-op *) T-op *) T-op *) T-op

SARHAR SAR Harness - *) storage *) T-op *) T-op *) T-op *) T-op

PDHAR Power and Data Handling Harness - *) storage *) T-op *) T-op *) T-op *) T-op

T-op = operat ing Temperature

(****) Any of the SSPA can however be used as operating unit

(***) For a non operating PHM , the Tnon-op temperature applies whatever the ion pump status is (on or of f )

PHM 1

INIT SequenceLaunch Pad to

Fairing Jet tisonFairing Jet t ison to

Separat ion

Stand By/Test M ode

on ground

1

PHM 2 1

Qty

Off M ode

storage&transport

(**) In order to minimise satellite power consumption while minimising the t ime to recover the mission (clocks remains on)

Tnon-op = non-operat ing Temperature

*): passive element

(*) all P/L modes are however possib le (except Launch mode)

S/C M ode

Equipment




Qty Temperature Qty Temperature Qty Temperature Qty Temperature Qty Temperature Qty Temperature

on To guarant ee on To guarantee on To guarantee on To guarantee on To guarantee on To guarantee

RAFS1 Rubidium Atomic Frequency Standard 1 1 0 Tnon-op 1 T-op(*) 1 T-op (*) 1 T-op(**) 0 Tnon-op 1 T-op (*)

RAFS2 Rubidium Atomic Frequency Standard 2 1 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op

Passive Hydrogen M aser 1 (***) 0 Tnon-op 1 T-op(*) 1 T-op (*) 1 T-op(**) 0 Tnon-op 1 T-op (*)

Ion pump (part of PHM 1) 1 Tnon-op 1 T-op (*) 1 T-op (*) 1 T-op(**) 1 Tnon-op 1 T-op (*)

Passive Hydrogen M aser 2 (***) 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op 0 Tnon-op

Ion pump (part of PHM 2) 1 Tnon-op 1 Tnon-op 1 Tnon-op 1 Tnon-op 1 Tnon-op 1 Tnon-op

CM CU Clock M onitoring and Cont rol Unit 1 0 Tnon-op 1 T-op(*) 1 T-op (*) 1 T-op(**) 0 Tnon-op 1 T-op (*)

PLSU Payload Security Unit 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

NSGU Navigat ion Signal Generator Unit 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

FGUU Frequency Generat ion and Upconversion Unit 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

IPTC1-3 Input Test Coupler 1-3 3 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

SPLIT1-3 Low Power HB Split ter 1-3 3 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

NAVHPA1 SSPA HB1 (L1)2 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

NAVHPA2 SSPA HB2 (L1)2 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

NAVHPA3 SSPA HB1/HB2 Spare (L1)2 1 0 Tnon-op 0 Tnon-op 0 Tnon-op (****) 0 Tnon-op 0 Tnon-op 0 Tnon-op (****)

NAVHPA4 SSPA LB1 (E5)2 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

NAVHPA5 SSPA LB1 Spare (E5)2 1 0 Tnon-op 0 Tnon-op 0 Tnon-op (****) 0 Tnon-op 0 Tnon-op 0 Tnon-op (****)

NAVHPA6 SSPA LB2 (E6)2 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

NAVHPA7 SSPA LB2 Spare (E6)2 1 0 Tnon-op 0 Tnon-op 0 Tnon-op (****) 0 Tnon-op 0 Tnon-op 0 Tnon-op (****)

NAVSW1-3 RF High Power Switch 1-3 3 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

NAVLOAD1-3 RF High Power Load 3 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

OM UX OM UX Diplexer LB (E5/E6) 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

OPF 1 Output Filter 1 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

OPF 2 Output Filter 2 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

OPTC1-3 Output Test Couplers 1-3 3 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

NAVANT L-Band Navigat ion Antenna 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

M ISREC C-Band Receiver 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

M ISANT C-band Antenna Array 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

SARANT SAR Rx/Tx Antenna 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

SARIPTC SAR Test Coupler 406 M Hz 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

SAROPTC SAR Test Coupler L-Band 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

SART SAR Transponder Assembly 1 0 Tnon-op 0 Tnon-op 1 T-op (*) 0 Tnon-op 0 Tnon-op 1 T-op (*)

RTU Remote Terminal Unit 1 0 Tnon-op 1 T-op (*) 1 T-op (*) 1 T-op 0 Tnon-op 1 T-op (*)

HPHAR01-17 RF Cable 1-17 17 *) T-op *) T-op *) T-op *) T-op *) T-op *) T-op

LPHAR RF LP Harness - *) T-op *) T-op *) T-op *) T-op *) T-op *) T-op

CBHAR C-Band Harness - *) T-op *) T-op *) T-op *) T-op *) T-op *) T-op

SARHA R SAR Harness - *) T-op *) T-op *) T-op *) T-op *) T-op *) T-op

PDHAR Power and Data Handling Harness - *) T-op *) T-op *) T-op *) T-op *) T-op *) T-op

(***) For a non operat ing PHM , the Tnon-op temperature applies whatever the ion pump status is (on or of f )

(****) Any of the SSPA can however be used as operat ing unit

T-op = operat ing Temperature

PHM 1

Tnon-op = non-operat ing Temperature

*): passive element

(*) all P/L modes are however possible (except Launch mode)

(**) In order to minimise satellite power consumpt ion while minimising the t ime to recover the mission (clocks remains on)

Sun Aquisit ion M ode

Equipment

1

PHM 2 1

Qty

S/C M ode Orbit Change M ode Safe M ode Intermediate Safe M odeEarth Aquisit ion

M ode Normal M ode

Table 20 - Payload unit's configuration




4.2.2.5 Satellite mode transition

ENVREQ-225 : Satellite mode transitions The satellite mode transitions as defined in the next table shall be considered in the thermal control design:

Goto From:

IDM

SAM EAM NM OCM ISM SM

IDM x SAM x (x) EAM x x (x) NM x x (x) (x) OCM x (x) ISM (x) (x) SM (x)

X: Nominal transition (X): In case of emergency only

Table 21 - Satellite mode transitions

ENVREQ-228 : Transition from SAM to NM: S/C Attitude: The satellite is in any attitude between SAM and NM configuration. The satellite attitude control provides that during transition to SAM and EAM, a S/C face (except -X, -Z and +Z) will not remain continuously sun illuminated during more than TBC (15 min.) time. Duration of transition:

• Earth pointing EAM is achieved within 30min (TBC) • The satellite attitude control shall provide that pitch and roll accuracy with stable

Earth pointing is achieved within 15 min. (TBC) after EAM acquisition. • The satellite attitude control shall provide that Yaw steering with stable Earth pointing

is achieved within 90 min. (TBC) after EAM acquisition. • Transition time from EAM to NM with full performance is achieved within maximum 90

min. (TBC).

ENVREQ-229 : Transition from NM to SM: S/C Attitude: The satellite attitude is any between NM and SM configuration The satellite attitude control provides that during transition to SM, a S/C face (except -X) will not remain continuously sun illuminated during more than 15 min. (TBC). Duration of transition: The satellite attitude control provides that the transition from NM to SM with stable sun pointing is achieved within 30min. (TBC), excluding eclipse time.




4.2.3 Ground Facilities Thermal Environments

ENVREQ-231 : Integration, transportation and testi ng The following ground phases shall be considered:

• Integration • Handling of S/C • System tests • Storage

ENVREQ-232 : Climatic conditions on ground The climatic conditions in the ground facilities and in the transport container shall be maintained as described in the table below

Location Temperature (°C) Relative Humidity (%)

Pressure (bar)

Container 17-30°C 0 – 60 % 0.7 – 1.1 bar

Clean rooms 20 +/- 3°C 55 +/-10 % Ambient

Table 22 - Climatic conditions on ground

ENVREQ-235 : Status of the propulsion system on gr ound For storage and transport the propulsion system is dry (without propellant).

ENVREQ-236 : Status of the PHM on ground The stay-alive power (ion pump) of each PHM clock needs to be on continuously after acceptance performance testing of the equipment.

ENVREQ-237 : Transportation The transportation equipment including container shall withstand the environment as specified below, while maintaining the unit it contains within its applicable transport temperature range:

Location Temperature (°C) Relative Humidity (%)

External -50/+80°C 0 –100 %

Table 23 - Transportation equipment environment




Note: Transportation by air shall be done using pressurised cargo bay. 1. Margin for pressure drop during air transportation TBD. 2. Rate of pressure change: for rate of pressure change 45 mbar/s (global) and 5

mbar/s (local) shall be assumed 3. Number of cycles of pressure change: max 12 per aircraft transportation 4. The transport container shall not be pressurised (“breathing container”)

4.2.4 Pre-Launch Phase As the satellite shall be compatible with Ariane 5, Proton, Soyuz and Zenit the envelope conditions of all candidate launchers have been established.

ENVREQ-246 : Pre-launch phase, on ground On ground before launch, the following phases shall be considered: 1. S/C preparation within S/C preparation building and transport between the buildings in the transport container 2. S/C mated on dispenser and mated to launch vehicle (LV) inside Final Assembly Building and encapsulated inside fairing 3. Transfer from Final Assembly Building to launch pad

4.2.5 Activities in S/C Preparation Building

ENVREQ-248 : Environment conditions in the integra tion/ preparation facilities and transport containers The environmental conditions of the integration- and preparation facilities and the transport containers which are used for transport between ground facilities and launch pad are defined in hereunder table and shall be considered.

Cleanliness (Class)

Temperature (°C) Relative Humidity (%)

S/C-Preparation Building 100 000 20°C ≤ 30°C 60% ±10%

Satellite Container 100 000 20°C ≤ 30°C ≤ 60% with GN2 flushing

Table 24 - Environment conditions in the integrat ion/ preparation facilities and transport containers




4.2.6 Spacecraft Configuration at Launch

4.2.6.1 General configuration

ENVREQ-253 : Satellite on launcher configurations The following S/C configurations within the launcher shall be considered, as depicted in the figure below: • Single launch: 1 S/C is fixed horizontally on the single launch adapter. • Multiple launch:

o 2 S/C are fixed on the multiple launch dispenser o 6 S/C are fixed on the multiple launch dispenser

XS/C Solar Array

Dispenser

ZS/C

YS/C

XS/C

Solar Array Solar Array

Dispenser

XS/C

ZS/CZS/C YS/C YS/C

Figure 8 Soyuz and Zenit launch configuration (sing le and dual launch)




XS/C


Dispenser

XS/C

ZS/CZS/C YS/C YS/C

ZS/C

ZS/C

Multiple Launch Configuration for 6 S/C for Ariane (TBC)

A

View A

ZS/C

ZS/C

ZS/C

ZS/C

Figure 9 Ariane 5 launch configuration




XS/C


Dispenser

XS/C

ZS/CZS/C YS/C YS/C

ZS/C

Multiple Launch Configuration for 6 S/C for Proton (TBC)

A

View A

ZS/C

ZS/C

Figure 10 Proton launch configuration




4.2.6.2 Thermal Interfaces with Launcher

4.2.6.2.1 Conductive Interface

ENVREQ-262 : Conductive interface satellite to lau ncher The S/C shall be thermally decoupled form the launcher dispenser. The conductive conductance of each separation point shall be smaller than 0,25 W/K (TBC). Four separation points are used for the fixation between S/C and dispenser.

4.2.6.2.2 Radiative Interface

ENVREQ-264 : Radiative interface satellite to laun cher The dispenser of the launcher shall be covered with MLI (TBC).

4.2.6.3 S/C Configuration within all S/C modes The S/C configuration and associated thermal environments during the flight missions are defined by the following S/C-modes:

• Launch Pad to Fairing Jettison • Fairing Jettison to Separation • INIT Sequence • Sun Acquisition • Earth Acquisition • Normal Mode • Orbit Change Mode • Safe Mode • Intermediate Safe Mode

The definition of the satellite units on/off status according to the platform and payload states for the different spacecraft modes is shown in the tables of sections 4.2.2.3 and 4.2.2.4.

4.2.7 Under Fairing Conditions before Launch

ENVREQ-267 : Under fairing conditions before launc h The fairing cavity is vented during transfer of the upper composite (regardless of whether it is integrated to the LV) and during the stand-by phase on the launch pad except during the erection of the LV on the pad. The air-conditioning characteristics which shall be considered are depicted in hereunder table. These conditions are given w/o considerations of the operational phase. The values are given for the selected launchers separately. Additional information which is missing, e.g. Air Velocity (except for Ariane) will be negotiated with the launcher authorities and will be provided in a future issue of the present document. For special operational steps the air flow and the temperatures can be agreed upon the user´s requirements with the launcher authorities.




Environmental Conditions

Cleanliness 100 000 Class

Envelope temperature for the injected air temperature

• Soyuz

• Zenit

• ARIANE 5

• Proton

18 to 30°C (Tolerance: -3/+5°C)

18° C ≤T≤ 25°C (Tolerance: 3/+10°C)

18° C ≤T≤ 30°C (Tolerance: -3/+5°C)

18° C ≤T≤ 25°C (Tolerance: -3/+10°C)

18° C ≤T≤ 30°C (Tolerance: -3/+5°C)

Relative humidity 5-60 %

Ventilation air flow adjustable

• Soyuz

• Zenit

• ARIANE 5

• Proton

≤ 6000 m³/h

≤ 10000 m³/h

4000 m³/h

13000 m³/h

Air velocity under fairing

• Soyuz

• Zenit

• ARIANE 5

• Proton

TBD

TBD

≤ 2 m/s with local variations possible

TBD

Filtration 0,3 µm

Table 25 - Under fairing conditions before launch

ENVREQ-272 : Functional status of the units before launch The on/off status of the satellite units in this phase is defined in the tables of sections 4.2.2.3 and 4.2.2.4 (column "Launch Pad to Fairing Jettison").

4.2.8 Time Line from Launch to SA Deployment Typical mission profiles are presented for the launch phase (Ariane, Proton, Soyuz, Zenit) and the on-orbit phase which starts with the separation of the S/C from the launcher. At present only an overview over the various options is given. A detailing will be performed in the phase C/D, based on the selection of launchers and based on the information of the launcher authorities on the availability of the launchers (e.g. type of Ariane 5) and the launch site (Soyuz). For preliminary layout activities of the TCS, the worst timeline shall be taken.




4.2.8.1 Soyuz launch For Galileo a launch from Kourou is foreseen. Detailed information up to now is only available for a launch from Baikonur. For the launch from Kourou the typical sequence of events will be similar. A significant difference will occur for the geographical location of the launcher ground track and of the point where the separation from the launcher takes place. In the following the sequence for launch from Baikonur is given.

• Lift-off from Baikonur • Fairing jettisoning about 187 sec after lift-off • Fregat separation onto 200 km circular parking orbit • Coast phase and first burn to reach the transfer orbit (t=1570 s) • (ha = 23616 km, hp = 200 km, i = 52.5° ) • Fregat engine extinction (t=2240 s) • Fregat coast phase up to apogee • Second Fregat burn for transfer to final orbit (t=14300 s) • Insertion into final orbit (t=14590 s) • Transition of Fregat to the desired three axis stabilized attitude for S/C separation

(launcher axis XL perpendicular to orbit plane, separation of S/C 1 in the flight direction, separation of S/C 2 anti-parallel to the flight direction, Φ = 0deg)

• Separation of both S/C (about 4 - 4.5 hours after lift-off) • Fregat transition to graveyard orbit (two manoeuvres)

4.2.8.2 Zenit launch Zenit with upper stage DM-SL (enhanced version)

• Lift-off from Baikonur • Fairing jettisoning about 320 sec after lift-off • Insertion of upper stage into initial orbit (ha = 255 km, hp = -2082 km) • Coast phase (about 10 s) and first burn to reach the intermediate orbit • Coast phase (about 3870 s) and second burn to reach the transfer orbit • Coast phase (about 11890 s) up to apogee • Third burn for transfer to final orbit • Transition of DM - SL to the desired three axis stabilized attitude for separation of the

two S/C in plane 1 (launcher axis XL perpendicular to orbit plane, Φ = 0 deg) • Separation of the two S/C in plane 1 • Transition of upper stage to the desired three axis stabilized attitude for the

separation of the two S/C in plane 2 (launcher axis perpendicular orbit plane, Φ = 60 deg)

• Separation of the two S/C in plane 2 (about 5 - 5.5 hours from lift-off) o Transition of upper stage to graveyard orbit (two manoeuvres)

4.2.8.3 Proton launch Lift-off from Baikonur

• Fairing jettisoning about 342 sec after lift-off • Insertion into parking orbit (first upper stage burn)




• Insertion into intermediate orbit (ha = 4700 km, hp = 240 km, i = 51.9°, second upper stage burn)

• Third upper stage burn (t=12420 s) • Jettison add-on tank • Insertion into transfer orbit by forth upper stage burn (t= 13260 s), • (ha = 23616 km, hp = 340 km, i = 52.5°) • Fifth upper stage burn (t=25080 s) • Insertion into final orbit (t = 25680 s) • Transition of upper stage to the desired three axis stabilized attitude for the

separation of three S/C in plane 1 (launcher axis XL perpendicular to orbit plane) • Separation of the three S/C in plane 1 • Transition of upper stage to the desired three axis stabilized attitude for the

separation of the three S/C in plane 2 (launcher axis XL perpendicular to orbit plane, rotation about XL to achieve new attitude)

• Separation of the three S/C in plane 2 (about 7.5 hours after lift-off) • Transition of upper stage to graveyard orbit (two manoeuvres)

4.2.8.4 AR5-ECA with Fregat upper stage This is the baseline solution for Ariane launch. As this version only is under investigation at present, only rough information is available. The ascent exists of the following main phases.

• Boost phase of EAP/EPC • Boost of ECA • First boost phase of Fregat • Coast phase • Second boost of Fregat • Transition of upper stage to the desired three axis stabilized attitude for the

separation of the three S/C in plane 1 (launcher axis XL perpendicular to orbit plane)

• Separation of the three S/C in plane 1 • Transition of upper stage to the desired three axis stabilized attitude for the

separation of the three S/C in plane 2 (launcher axis XL perpendicular to orbit plane, rotation about xL to achieve new attitude)

• Separation of the here S/C in plane 2 • Transition to graveyard orbit (two manoeuvres)

4.2.9 Depressurisation under fairing

ENVREQ-279 : Lift-off phase The following venting profile shall be taken. It is derived from Zenit and covers all other launchers:




Figure 11 - off phase venting profile




4.2.10 Thermal conditions under fairing

ENVREQ-283 : Satellite Phase: Launch Phase S/C Mode: From Launch pad until fairing jettison. Flight case: 'From Launch pad until fairing jettison': A maximum thermal flux density radiated by the fairing of 1000 W/m\'b2 at any point shall be taken. This value is derived from Ariane 5 and covers also Soyuz and Proton. For Zenit, no information is available at present. The maximum duration of this phase to consider is t= 342 sec. (TBC) The on/off status of the satellite units in this phase is defined in the tables of sections 4.2.2.3 and 4.2.2.4 (column: "Launch Pad to Fairing Jettison").

XS/C Solar Array

Dispenser

ZS/C

YS/C

Thermal Heatflux radiated byfairing to S/C

XS/C


Dispenser

XS/C

ZS/CZS/C YS/C YS/C

Fairing

Thermal Heatflux radiated byfairing to S/C

Figure 12 Thermal conditions under fairing

4.2.10.1 Thermal Conditions at fairing jettison

ENVREQ-287 : Flight case: 'Aerodynamic heating aft er fairing jettison': The following conditions shall be considered: The typical aero-thermal flux varies from the values defined in 1135W/m2 to less than 200 W/m2 within 10 seconds after the fairing jettisoning (TBC) as shown in the figure below. This value shall be taken for preliminary development. The final value which shall be applied will be specified together with the applicable launcher authorities. The aero-thermal flux applies normally on the +X side for multiple launch and on the +Z side for single launch The duration to consider for this phase is: t= 20 sec. (TBC) The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Launch Pad to Fairing Jettison").




XS/C Solar Array

Dispenser

ZS/C

YS/C

Aerodynamic Heating

XS/C


Dispenser

XS/C

ZS/CZS/C YS/C YS/C

Aerodynamic Heating

Figure 13 Aerodynamic heating after fairing jettiso n

4.2.10.2 Thermal conditions after fairing jettison until separation The following thermal conditions apply after fairing jettison up to separation:

ENVREQ-290 : 1. After fairing jettison: Cold Case for single and multiple launch: This orbit case is during transfer orbit during which the launcher may perform manoeuvres.




The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Fairing Jettison to Separation"). S/C configuration:

• S/C fixed to dispenser • Solar array in stowed position • The S/C is in shadow (no sun) for a maximum duration of tmax.= 7 hours.

• ENVREQ-291 : 2. Coast Phase Hot Case for multiple launch: • This orbit case is during transfer orbit where the launcher is not firing. • The on/off status of the satellite units in this phase is defined in the tables of section

4.2.2.3 and 4.2.2.4 (column S/C mode: "Fairing Jettison to Separation"). • For the launcher the following requirements apply: • Barbecue mode. • spin around S/C X-axis • spin rate min. 1°/sec., max. 3°/sec., Tolerance = 0,2°/sec. (TBC) • S/C configuration: • S/C fixed to the dispenser • Solar array in stowed position • S/C attitude: The angle between Orbiter /Spacecraft roll axis and Sun shall be: α =

60° with a variation range of: +/-20°. • Duration in this orientation: t = without duration limit = steady state




X


Dispenser

X

ZZ Y Y

0°

90°

a

60°

-20°

+20°

Launcher Roll Axis

spinSun

Figure 14 Coast Phase Hot Case for multiple launch (Angle between Orbiter /Spacecraft roll axis and Sun = 60°)

ENVREQ-294 : 3. Coast Phase Hot Case for multiple launch. This orbit case is during transfer orbit where the launcher is not firing. The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Fairing Jettison to Separation"). For the launcher the following requirements apply:

• Barbecue mode. • spin around S/C X-axis




• spin rate min. 1°/sec., max. 3°/sec., Tolerance =0 ,2°/sec. (TBC) S/C configuration:

• S/C fixed to the dispenser • Solar array in stowed position • S/C attitude: The angle between Orbiter /Spacecraft roll axis and Sun shall be: α = 0°. • Duration in this orientation: t = without duration limit = steady state

ENVREQ-295 : 4. Coast Phase Hot Case for multiple launch. This orbit case is during transfer orbit where the launcher is not firing. The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Fairing Jettison to Separation"). For the launcher the following requirements apply:

• Barbecue mode. • spin around S/C X-axis • spin rate min. 1°/sec., max. 3°/sec., Tolerance =0 ,2°/sec. (TBC)

S/C configuration:

• S/C fixed to dispenser • Solar array in stowed position • S/C attitude: The angle between Orbiter /Spacecraft roll axis and Sun shall be: α =

90°. • Duration in this orientation: t = without duration limit = steady state

ENVREQ-296 : 5. Coast Phase Hot Case for single la unch: This orbit case is during transfer orbit where the launcher is not firing. The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Fairing Jettison to Separation"). For the launcher the following requirements apply:

• Barbecue mode. • spin around S/C Z-axis • spin rate min. 1°/sec., max. 3°/sec., Tolerance =0 ,2°/sec. (TBC)

S/C configuration:


60° with a variation range of: +/-20°. • Duration in this orientation: t = without duration limit = steady state




XS/C Solar Array

Dispenser

ZS/C

YS/C

Launcher Roll Axis

spin

0°

90°

60°

Sun

Figure 15 Coast Phase Hot Case for single launch (A ngle between Orbiter /Spacecraft roll axis

and Sun = 60°)



S/C configuration:

• S/C fixed to dispenser • Solar array in stowed position • S/C attitude: The angle between Orbiter /Spacecraft roll axis and Sun shall be: α = 0°. • Duration in this orientation: t = without duration limit = steady state




XS/C Solar Array

Dispenser

ZS/C

YS/C

Launcher Roll Axis

spin

0°

90°

60°

Sun


and Sun = 0°)



S/C configuration:


90°. • Duration in this orientation: t = without duration limit = steady state




XS/C Solar Array

Dispenser

ZS/C

YS/C

Launcher Roll Axis

spin

0°

90°

60°

Sun


and Sun = 90°)

ENVREQ-305 : Conductive heat transfer through the launcher separation system For conductive heat transfer through the launcher separation system, the conductance shall be taken as specified in 4.2.6.2.1 and the interface temperature of the launcher interface (launcher side) shall be taken as TBD.

4.2.10.3 Orbit case: 'Sun Acquisition Mode ' The following conditions apply in Sun Acquisition Mode:

ENVREQ-307 : 1. Before Solar Generator deployment: The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Sun Acquisition"). S/C configuration:

• S/C separated from the dispenser. • Solar array is in stowed position.

S/C attitude: • -X axis is directly oriented to the sun. • S/C altitude: orbit height 23.222km. (semi major axis = 29.600,3km) • Duration in this orientation: t = without duration limit = steady state




Figure 18 Sun Acquisition Mode conditions (Before S olar Generator deployment)

ENVREQ-310 : 2. After Solar Generator deployment: The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Sun Acquisition"). S/C configuration:

• S/C separated from the dispenser. • Solar array is in deployed position.

S/C attitude: • -X axis is directly oriented to sun. • S/C altitude: orbit height 23.222km. (semi major axis = 29.600,3km) • Duration in this orientation: t = without duration limit = steady state

Figure 19 Sun Acquisition Mode conditions (After So lar Generator deployment)

4.2.10.4 Orbit case: 'Earth Acquisition Mode ' The following thermal conditions apply in Earth Acquisition Mode:

ENVREQ-314 : Earth Acquisition Mode Case 1: The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Earth Acquisition"). S/C configuration:




• S/C separated from the dispenser. • Solar array in deployed position.

S/C attitude: • +Z axis is directly oriented to earth. • +X axis in flight direction • S/C altitude: orbit height 23.222km. (semi major axis = 29.600,3km) • S/C orbit inclination 56° • The maximum sun elevation above the orbit plane is +/- 79,4° • Time: orbit period t = 14,08 hours (transient run)



S/C attitude: • +Z axis is directly oriented to earth. • +X axis in flight direction • S/C altitude: orbit height 23.222km. (semi major axis = 29.600,3km) • S/C orbit inclination 56° • The minimum sun elevation is 0° (maximum eclipse d uration of 59,6 min.) • Time: orbit period t = 14,08 hours (transient run)



S/C attitude: • +Z axis is directly oriented to earth. • +Y axis in flight direction • S/C altitude: orbit height 23.222km. (semi major axis = 29.600,3km) • S/C orbit inclination 56° • The maximum sun elevation above the orbit plane is + 79,4° or - 79,4° (no eclipse) • Time: orbit period t = 14,08 hours (transient run)



S/C attitude:




• +Z axis is directly oriented to earth. • +Y axis in flight direction • S/C altitude: orbit height 23.222km. (semi major axis = 29.600,3km) • S/C orbit inclination 56° • The minimum sun elevation is 0° (maximum eclipse d uration of 59,6 min.). • Time: orbit period t = 14,08 hours (transient run)



S/C attitude: • +Z axis is directly oriented to earth. • Yaw steering (rotation around Z-axis) • S/C altitude: orbit height 23.222km. (semi major axis = 29.600,3km) • S/C orbit inclination 56° • The maximum sun elevation above orbits plane is +79,4° or -79,4° (no eclipse) • Time: orbit period t = 14,08 hours (transient run)



S/C attitude: • +Z axis is directly oriented to earth. • Yaw steering (rotation around Z-axis) • S/C altitude: orbit height 23.222km. (semi major axis = 29.600,3km) • S/C orbit inclination 56° • The minimum sun elevation is 0° (maximum eclipse d uration of 59,6 min.). • Time: orbit period t = 14,08 hours (transient run)

4.2.10.5 Orbit case: 'Normal Mode ' The following thermal conditions apply in Normal Mode:

ENVREQ-321 : Normal Mode Case 1: The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Normal Mode"). S/C configuration:





S/C attitude: • +Z axis is directly oriented to earth. • Yaw steering (rotation around Z-axis) • S/C altitude: orbit height 23.222km. (semi major axis = 29.600,3km) • S/C orbit inclination 56° • The maximum sun elevation above the orbit plane is + 79,4° or - 79,4° (no eclipse) • Time: orbit period t = 14,08 hours (transient run)

ENVREQ-322 : Normal Mode Case 2: The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Normal Mode"). S/C configuration:



4.2.10.6 Orbit case: 'Orbit Change Mode ' The following thermal conditions apply in Orbit Change Mode:

ENVREQ-324 : Orbit Change Mode Case 1: This orbit case occurs in the orbit change mode for in plane manoeuvres: The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Orbit Change Mode"). S/C configuration:


S/C attitude: • +Z axis is directly oriented to earth. • +X axis in flight direction • S/C altitude: orbit height 23.222km. (semi major axis = 29.600,3km) • S/C orbit inclination 56° • The maximum sun elevation above the orbit plane is + 79,4° or - 79,4° (no eclipse) • Time: orbit period t = 14,08 hours (transient run)

ENVREQ-325 : Orbit Change Mode Case 2: This orbit case occurs in the orbit change mode for in plane manoeuvres: The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Orbit Change Mode").




S/C configuration: • S/C separated from the dispenser. • Solar array in deployed position.

S/C attitude: • +Z axis is directly oriented to earth. • +X axis in flight direction • S/C altitude: orbit height 23.222km. (semi major axis = 29.600,3km) • S/C orbit inclination 56° • The minimum sun elevation is 0° (maximum eclipse d uration of 59,6 min.). • Time: orbit period t = 14,08 hours (transient run)

ENVREQ-326 : Orbit Change Mode Case 3: This orbit case occurs in the orbit change mode for out-of-plane manoeuvres: The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Orbit Change Mode"). S/C configuration:


S/C attitude: • +Z axis is directly oriented to earth. • +Y axis in flight direction • S/C altitude: orbit height 23.222km. (semi major axis = 29.600,3km) • S/C orbit inclination 56° • The maximum sun elevation above orbits plane is + 79,4° or - 79,4° (no eclipse) • Time: orbit period t = 14,08 hours (transient run)

ENVREQ-327 : Orbit Change Mode Case 4: This orbit case occurs in the orbit change mode for out-of-plane manoeuvres: The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Orbit Change Mode"). S/C configuration:


S/C attitude: • +Z axis is directly oriented to earth. • +Y axis in flight direction • S/C altitude: orbit height 23.222km. (semi major axis = 29.600,3km) • S/C orbit inclination 56° • The minimum sun elevation is 0° (maximum eclipse d uration of 59,6 min.) • Time: orbit period t = 14,08 hours (transient run)

4.2.11 Orbit case: 'Safe Mode ' The following thermal conditions apply in (Ultimate) Safe Mode:




ENVREQ-329 : Safe Mode The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Safe Mode"). S/C configuration:


S/C attitude: • -X axis is directly oriented to sun. • S/C altitude: orbit height 23.222km. (semi major axis = 29.600,3km) • Time in this position t = without duration limit = steady state.

4.2.12 Orbit case: 'Intermediate Safe Mode ' The following thermal conditions apply in Intermediate Safe Mode:

ENVREQ-331 : Intermediate Safe Mode Case 1: This mode is used as autonomous recovering mode (after on-board failure), providing earth pointing with relaxed pointing performance. The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Intermediate Safe Mode"). S/C configuration:


S/C attitude: • +Z axis is directly oriented to earth. • Yaw steering (rotation around Z-axis) • S/C altitude: orbit height 23.222km. (semi major axis = 29.600,3km) • S/C orbit inclination 56° • The maximum sun elevation above the orbit plane is + 79,4° or - 79,4° (no eclipse) • Time: orbit period t = 14,08 hours (transient run)

ENVREQ-332 : Intermediate Safe Mode Case 2: This mode is used as autonomous recovering mode, providing earth pointing with relaxed pointing performance. The on/off status of the satellite units in this phase is defined in the tables of section 4.2.2.3 and 4.2.2.4 (column S/C mode: "Intermediate Safe Mode"). S/C configuration:






4.2.12.1 Yaw Steering Law This section defines the yaw steering law which fulfils the following objectives (cf. figure below): • Maintain the sun direction within the (X, Z) spacecraft plane except when the sun

direction is close to the spacecraft Z axis (for low sun elevation onto the orbit plane and around local noon and midnight) where the sun elevation onto the (X, Z) plane is limited to 2 deg (in order to allow the solar array to be pointed to the sun via the one axis SADM),

• Maintain the sun within the -X half space in order to avoid any solar flux within the +X (clock) radiator.

ψ = βx0

x

y

z

6 PM

ψ = 180° − β

x0

x

y

z6 AM

σ = −90°

N

E

SUNAngle

MIDNIGHT

y

z

x0SADM - Angle

σ = −β

ψ = 90°

x

NOON

y

Z

x0

SADM - Angle

σ = −(180° − β)

ψ = 90° x

σ = −90°

η

β

Figure 20 Yaw steering law

ENVREQ-337 : Definition of nominal attitude The spacecraft nominal attitude applying for yaw steering law is defined as follows: • The satellite +Z axis is directed to the earth • The yaw attitude (around the Z axis) follows the yaw steering law defined as follows:




ksXr .)cos( =ψ (a)

ksYr .)sin( =ψ (b)

21

1

Zsk

−−= (c)

, sin( ) sin( )

,H OX X OY Y

O

s s and ss

s elsewhere

β β < <=

(d)

βX = 15 deg., βY = 2 deg. (e) 2 2, , 1 ( )T

H OX HY OX HY OZs s s s s sign s = − − ⋅

(f)

( )sin( ) sin( )sin( )

OXHY Y Y OY

X

ss sigy sigy sβ β

β= ⋅ − ⋅ − ⋅

(g)

Where: ψr : yaw reference angle (w.r.t the orbital reference system) sO

T = (sOX , sOY , sOZ )

T: sun reference vector (in the orbital reference system) sH

T = (sHX , sHY , sHZ )

T: sun auxiliary vector, as defined above (used in the auxiliary region where the sun vector is close to the Z axis)

sT = (sX , sY , sZ )

T: pseudo sun reference vector, as defined above orbital reference system (Xo, Yo, Zo): + Zo pointed to the Earth, + Xo along the satellite velocity, Yo completes the directly oriented axes reference system βX / βY : angles defining the auxiliary region where sO is substituted by sH sigy : sign of sOY at the beginning of the auxiliary region (i.e. at the time where sOX = ±sin(βX) )

ENVREQ-339 : Accuracy in NM mode: With respect to the nominal attitude in normal mode as defined before, the following satellite attitude errors shall be taken into account in the thermal analyses, where relevant:

• Roll/Pitch error: ± 0.3 deg • Yaw error: ± 0.8 deg

4.2.13 Thermal Space environment

4.2.13.1 Space Environment The relevant space thermal environment includes:

• Solar and Planetary radiation • Space sink • Lifetime




4.2.13.1.1 Definition of Solar, Albedo, and Earth fluxes

ENVREQ-343 : Solar Constant The following values for Solar constant shall be used: The solar constant for in orbit condition is: • mean value at 1 AU = 1327 ± 5 W/m2 • perihelion = 1418 W/m2 • aphelion = 1326 W/m2.

ENVREQ-344 : Earth Albedo and Earth flux The following values for earth albedo and earth flux shall be taken: • Earth Albedo: The average Earth Albedo is 0.3. • Earth flux: The infrared power radiated by the Earth is 237 W/m2.

4.2.13.1.2 Solar Radiation

ENVREQ-346 : Total solar irradiance For the total solar irradiance under consideration of changes in earth/sun distance during the year the values of the figure below table shall be taken:

Figure 21 Total solar irradiance




4.2.13.1.3 Solar Spectral Irradiance

ENVREQ-350 : Solar spectral irradiance For computations which require solar irradiation data over narrow wave length bands, the solar spectral irradiance values as given in the following table shall be taken. The estimated error in these values is ± 5% in the wave length range of 0.3 to 3.0 µm. All values are for a mean earth/sun distance of 1.496x108 km. In the 0.3 to 0.75 µm range, the value Pλ is the average irradiance for a 10 nm band centred at the corresponding wave length.

Table 26 - Solar spectral irradiance




4.2.13.1.4 Earth Albedo and Emitted Radiation

ENVREQ-353 : Earth Albedo and Emitted Radiation The fraction of solar radiation reflected (Albedo) and the radiation emitted by Earth shall be considered as well as direct solar radiation for the design of the satellite and its externally mounted units, if relevant. Albedo and earth emitted radiation shall be considered in the thermal design and selection of operating characteristics of optical sensor systems.

4.2.13.1.5 Earth Radiation

ENVREQ-355 : Earth Radiation Between 8 and 12 µm, the thermal spectral radiation emitted by the earth atmosphere shall be considered as the radiation emitted by a black body at 254 K. For longer wave length, it shall be considered as the radiation emitted by a black body at 218 K.

4.2.13.1.6 Earth Albedo Radiation

ENVREQ-357 : Earth Albedo Radiation The global earth albedo is the ratio of the total radiation reflected from the earth to the solar incident of the earth. The spectral distribution shall be considered as approximately equivalent to the one emitted by a black body at 5760 K.

4.2.13.1.7 Space Sink Temperature

ENVREQ-359 : Space Sink Temperature The radiation temperature of cold space shall be considered as 4 K.

4.2.14 Thermal in orbit environment

ENVREQ-361 : Thermal in orbit environment The Satellite thermal control shall be sized to withstand the worst thermal environment loads due to the GALILEO orbital configuration.

4.2.14.1 Orbit definition and resulting sun illumi nation

ENVREQ-362 : Orbit type The orbit type for the constellation to be considered shall be a Walker 27/3/1 orbit, i.e. 9 operational spacecraft in each of 3 orbital planes, with the ascending nodes of the orbital planes uniformly distributed round the equator at intervals of 120 degrees.

ENVREQ-363 : Orbit and attitude characteristics The orbit and attitude characteristics are:




• Semi Major Axis: 29 600.3 km (from earth center) • Inclination: 56° • Eccentricity: 0 (+/- eps)

The sun elevation above orbit plane varies between: [-79.5°; and +79.5°] This orbit data results in the following: Period T=50682 s (14.08h) 7.1% max eclipse ratio (0.993 h) Note: Moon eclipse occurrence and duration (as well as potential cumulation to Earth eclipse) is 80 min (TBC). The resulting sun-angle elevation of the orbit is shown below:

Variation of the sun illumination angle ββββ' during the year

0.00

15.00

30.00

45.00

60.00

75.00

90.00

0 50 100 150 200 250 300 350

Day of the year

Angl

e de

gree

]

Figure 22 - angle elevation of the orbit

4.2.15 Unit Temperature Limits The following sections present the temperature requirements of all satellite units.

ENVREQ-948 : Unit reliability calculation For the unit reliability calculation the reliability temperatures (T rel.) which shall be used are presented in the tables of § 4.2.15.1 (T rel.).




4.2.15.1 Temperature limits of Platform Units

4.2.15.1.1 Temperature limits of TT&C Units

ENVREQ-368 : TX/RX (S-Band Transponder):

Unit: TX/RX (S-BAND TRANSPONDER) T min Tmax

Qualification -25 65 Acceptance -20 60 Operating temperature (°C)

TCS Design -15 55

Qualification -45 70 Acceptance -40 65 Non-Operating temperature (°C)

TCS Design -35 60 Start-up temperature (°C) -25 None

Ground storage and transport temp. (°C) -20 70

T rel. (°C) 30

Stability requirement /short term None

Stability requirement /long term None

Temperature limit definition (applicability) Base-plate (TRP)

Table 27 - Temperature limits of TX/RX (S-Band Tr ansponder)

ENVREQ-371 : TTCHYB (Hybrid):

Unit: TTCHYB (Hybrid) T min Tmax


TCS Design -35 55


TCS Design -40 60

Start-up temperature (°C) -45 None


T rel. (°C) 30




Table 28 - Temperature limits of TX/RX (TTCHYB (H ybrid))




ENVREQ-374 : TTCANTN and TTCANTZ (S-Band Antenna N adir/Zenith):

Unit: TTCANTN and TTCANTZ (S-Band Antenna Nadir/Zenith)

T min Tmax


Design -100 105


TCS Design -135 130





Temperature limit definition (applicability) Radome

Table 29 - Temperature limits of TTCANTN and TTCA NTZ (S-Band Antenna Nadir/Zenith)

ENVREQ-377 : TTCHAR (TTC RF-Harness):

Unit: TTCHAR (TTC RF-Harness): T min Tmax


TCS Design -100 105


TCS Design -185 120





Temperature limit definition (applicability) NA

Table 30 - Temperature limits of TTCHAR (TTC RF-H arness)




ENVREQ-380 : LRR (Laser Retro Reflector):

Unit: LRR (Laser Retro Reflector) T min Tmax


TCS Design -145 125


TCS Design -145 125

Start-up temperature (°C) N.A. None





Table 31 - Temperature limits of LRR (Laser Retro Reflector)

4.2.15.1.2 Temperature limits of PFSU Units

ENVREQ-384 : PFSU (Platform Security unit):

Unit: PFSU (Platform Security Unit) T min Tmax


TCS Design -25 50


TCS Design -30 65



T rel. (°C) 35




Table 32 - Temperature limits of LRR (Laser Retro Reflector)

4.2.15.1.3 Temperature limits of EPS Units




ENVREQ-388 : PCDU (Power Controlling and Distribut ion Unit)

Unit: PCDU (Power Controlling and Distribution Unit

T min Tmax


TCS Design -20 50


TCS Design -25 55



T rel. (°C) 35




ENVREQ-390 : SA-Y & SA+Y (Solar Array -Y & +Y)

Unit: SA-Y & SA+Y (Solar Array –Y & +Y) T min Tmax


TCS Design -170 120

Qualification -180 130

Acceptance -175 125 Non-Operating temperature (°C)

TCS Design -170 120


Ground storage and transport temp. (°C) -20 90 T rel. (°C) To be justified by SA supplier



Temperature limit definition (applicability) panels

Table 33 - Temperature limits of SA-Y & SA+Y (Sol ar Array -Y & +Y)




ENVREQ-393 : SADM-Y & SADM+Y (Solar Array Drive Me chanism -Y & +Y):

Unit: SADM-Y & SADM+Y (Solar Array Drive Mechanism –Y & +Y)

T min Tmax


TCS Design -15 60


TCS Design -40 75



T rel. (°C) 40



Temperature limit definition (applicability) Housing

Table 34 - Temperature limits of SADM-Y & SADM+Y (Solar Array Drive Mechanism -Y & +Y)

ENVREQ-396 : SAHD-Y1,2,3,4 & SAHD+Y1,2,3,4 (Solar Array Hold-down -Y & +Y):

Unit: SAHD-Y1,2,3,4 & SAHD+Y1,2,3,4 (Solar Array Hold-down –Y & +Y)

T min Tmax


TCS Design -30 65


TCS Design -30 65





Temperature limit definition (applicability) Base-plate (TRP) of bracket

Table 35 - Temperature limits of SAHD-Y1,2,3,4 & SAHD+Y1,2,3,4 (Solar Array Hold-down -Y & +Y)




ENVREQ-399 : BATT (Battery):

Unit: BATT (Battery) T min Tmax

Qualification 5 40 Acceptance 10 35 Operating temperature (°C)

TCS Design 15 30

Qualification 5 40 Acceptance 10 35 Non-Operating temperature (°C)

TCS Design 15 30

Start-up temperature (°C) 5 None Ground storage and transport temp. (°C) -5 35

T rel. (°C) 25




Table 36 - Temperature limits of BATT (Battery)

4.2.15.1.4 Temperature limits of Avionic Units

ENVREQ-403 : ICDU (Integrated control and data han dling unit):

Unit: ICDU (Integrated control and data handling unit)

T min Tmax


TCS Design -25 50


TCS Design -30 55



T rel. (°C) 30




Table 37 - Temperature limits of ICDU (Integrated control and data handling unit)




ENVREQ-406 : FSS1 & FSS2 (Fine Sun Sensor):

Unit: FSS1 & FSS2 (FSS, Fine Sun Sensor)

T min Tmax


TCS Design -25 60


TCS Design -30 65



T rel. (°C) 30




Table 38 - Temperature limits of FSS1 & FSS2 (Fi ne Sun Sensor)

ENVREQ-409 : CSS1 & CSS2 (Coarse Sun Sensors):

Unit: CSS1 & CSS2 (Coarse Sun Sensors)

T min Tmax


TCS Design -90 110


TCS Design -90 110

Start-up temperature (°C) None

(passive unit) None


T rel. (°C) 50




Table 39 - Temperature limits of CSS1 & CSS2 (Co arse Sun Sensors)




ENVREQ-412 : ES1 & ES2 (Earth Sensor 1 & 2):

Unit: ES1 & ES2 (Earth Sensor 1 & 2): T min Tmax


TCS Design -20 45


TCS Design -20 45

Start-up temperature (°C) -25 None Ground storage and transport temp. (°C) -20 55

T rel. (°C) 30



Temperature limit definition (applicability) Electrical Housing

Table 40 - Temperature limits of ES1 & ES2 (Earth Sensor 1 & 2)

ENVREQ-415 : GYRO (Gyro):

Unit: GYRO (Gyro) T min Tmax


TCS Design -25 60


TCS Design -30 65


T rel. (°C) 40




Table 41 - Temperature limits of GYRO (Gyro)




ENVREQ-418 : MTR1 & MTR2 (Magnetic Torquer 1 & 2):

Unit: MTR1 & MTR2 (Magnetic Torquer 1 & 2) T min Tmax


TCS Design -25 62


TCS TCS Design -30 70


T rel. (°C) 30



Temperature limit definition (applicability) Bar

Table 42 - Temperature limits of MTR1 & MTR2 (Mag netic Torquer 1 & 2)

ENVREQ-421 : RW1 & RW2 & RW3 & RW4 (Reaction Wheel 1, 2, 3, 4):

Unit: RW1 & RW2 & RW3 & RW 4 (Reaction Wheel 1,2,3,4)

T min Tmax


TCS Design 0 45


TCS Design -10 45



T rel (°C) 30




Table 43 - Temperature limits of RW1 & RW2 & RW3 & RW4 (Reaction Wheel 1, 2, 3, 4)




4.2.15.1.5 Temperature limits of Propulsion Units

ENVREQ-425 : PROTANK (Tank):

Unit: PROTank (Tank) T min Tmax


TCS Design 12 50


TCS Design 12 50

Start-up temperature (°C) N.A. None Ground storage and transport temp. (°C)

Dry configuration = without fuel (N 2H4) -20 60



Temperature limit definition (applicability) Tank

Table 44 - Temperature limits of PROTANK (Tank)

ENVREQ-428 : PROFVV (Fill and Vent Valve):

Unit: PROFVV (Fill and Vent Valve) T min Tmax


TCS Design 12 50


TCS Design 12 50






Table 45 - Temperature limits of PROFVV (Fill and Vent Valve)




ENVREQ-431 : PROPT (Pressure Transducer):

Unit: PROPT (Pressure Transducer) T min Tmax


TCS Design 12 50


TCS Design 12 50



T rel. (°C) 30




Table 46 - Temperature limits of PROPT (Pressure Transducer)

ENVREQ-434 : PROLF (Filter):

Unit: PROLF (Filter) T min Tmax


TCS Design 12 50


TCS Design 12 50






Table 47 - Temperature limits of PROLF (Filter)




ENVREQ-437 : PROFDV (Fill and Drain Valve):

Unit: PROFDV (Fill and Drain Valve) T min Tmax


TCS Design 12 50


TCS Design 12 50






Table 48 - Temperature limits of PROFDV (Fill and Drain Valve)

ENVREQ-440 : PROLVA & PROLVB (Latching Valve 1 & 2 ):

Unit: PROLVA & PROLVA (Latching Valve 1 & 2)

T min Tmax


TCS Design 12 50


TCS Design 12 50


Ground storage and transport temp. (°C) Dry configuration = without fuel (N 2H4)

-20 60




Table 49 - Temperature limits of PROLVA & PROLVB (Latching Valve 1 & 2)




ENVREQ-443 : PROTPA & PROTPB (Test Port A&B):

Unit: PROTPA & PROTPB (Test Port A&B) T min Tmax


TCS Design 12 50


TCS Design 12 50






Table 50 - Temperature limits of PROTPA & PROTPB (Test Port A&B)

ENVREQ-446 : PIPS (Piping):

Unit: PIPS (Piping) T min Tmax


TCS Design 12 50


TCS Design 12 50





Temperature limit definition (applicability) On Pipes

Table 51 - Temperature limits of PIPS (Piping)




ENVREQ-449 : PRORCT1A & PRORCT2A & PRORCT3A & PROR CT4A & PRORCT1B & PRORCT2B & PRORCT1B & PRORCT1B & (Reaction Control Thruster 1A,2A,3A,4A, 1B,2B,3B,4B):

Unit: PRORCT1A & PRORCT2A & PRORCT3A & PRORCT4A & PRORCT1B & PRORCT2B &

PRORCT1B & PRORCT1B & (Reaction Control Thruster 1A,2A,3A,4A, 1B,2B,3B,4B):

T min Tmax

Qualification 4 70 Acceptance 7 65

Operating temperature (°C): continuous firing

TCS Design 12 60

Qualification 4 90

Acceptance 7 85 Transient temperature during soak

back TCS Design 12 80


TCS Design 12 50





Temperature limit definition (applicability) Electro-valve

Table 52 - Temperature limits of PRORCT1A & PRORC T2A & PRORCT3A & PRORCT4A & PRORCT1B & PRORCT2B & PRORCT1B & PRORCT1B & (Reacti on Control Thruster

1A,2A,3A,4A, 1B,2B,3B,4B)

4.2.15.2 Temperature limits of Payload Units

ENVREQ-453 : RAFS1 & RAFS2 (Rubidium Atomic Freque ncy Standard 1 & 2):




Unit: RAFS1 & RAFS2 (Rubidium Atomic Frequency Standard 1 & 2)

T min Tmax


TCS Design -5 10


TCS Design -15 60


Ground storage and transport temp. (°C) - 10 50

T rel. (°C) 10

Stability requirement /short term +/- 1°C per 24 hours



Table 53 - Temperature limits of RAFS1 & RAFS2 (R ubidium Atomic Frequency Standard 1 & 2)

ENVREQ-456 : PHM1 & PHM2 (Passive Hydrogen Maser 1 & 2):

Unit: PHM1 & PHM2 (Passive Hydrogen Maser 1 & 2):

T min Tmax


TCS Design -5 10


TCS Design -15 40


Ground storage and transport temp. (°C) - 10 50

T rel. (°C) 10




Table 54 - Temperature limits of PHM1 & PHM2 (Pas sive Hydrogen Maser 1 & 2)




ENVREQ-459 : CMCU (Clock Monitoring and Control Un it):

Unit: CMCU (Clock Monitoring and Control Unit)

T min Tmax


TCS Design -15 40


TCS Design -40 70



T rel. (°C) 25




Table 55 - Temperature limits of CMCU (Clock Moni toring and Control Unit)

ENVREQ-462 : PLSU (Payload Security Unit):

Unit: PLSU (Payload Security Unit): T min Tmax


TCS Design -20 55


TCS Design -35 65


T rel. (°C) 35




Table 56 - Temperature limits of PLSU (Payload Se curity Unit)




ENVREQ-465 : NSGU (Navigation Signal Generator Uni t):

Unit: NSGU (Navigation Signal Generator Unit)

T min Tmax


TCS Design -10 40


TCS Design -35 70



T rel (°C) 30




Table 57 - Temperature limits of PLSU (Payload Se curity Unit)

ENVREQ-468 : FGUU (Frequency Generation and Up-con version Unit):

Unit: FGUU (Frequency Generation and Up-conversion Unit)

T min Tmax


TCS Design -10 40


TCS Design -40 70



T rel. (°C) 30




Table 58 - Temperature limits of FGUU (Frequency Generation and Up-conversion Unit)




ENVREQ-480 : IPTC1 & IPTC2 & IPTC3 (Input Test Cou pler):

Unit: IPTC1 & IPTC2 & IPTC3 (Input Test Coupler)

T min Tmax


TCS Design -40 85


TCS Design -55 100






Table 59 - Temperature limits of IPTC1 & IPTC2 & IPTC3 (Input Test Coupler)

ENVREQ-483 : SPLIT1 & SPLIT2 & SPLIT3 (Low Power S plitter):

Unit: SPLIT1 & SPLIT2 & SPLIT3 (Low Power Splitter)

T min Tmax


TCS Design -40 85


TCS Design -55 100






Table 60 - Temperature limits of SPLIT1 & SPLIT2 & SPLIT3 (Low Power Splitter)




ENVREQ-486 : NAVHPA1 & NAVHPA2 & NAVHPA3 & NAVHPA4 & NAVHPA5 & NAVHPA6 & NAVHPA7 (Solid State Power Amplifier):

Unit: NAVHPA1 & NAVHPA2 & NAVHPA3 & NAVHPA4 & NAVHPA5 & NAVHPA6 & NAVHPA7

(Solid State Power Amplifier ) T min Tmax


TCS Design -5 40


TCS Design -35 80



T rel. (°C) 30




Table 61 - Temperature limits of NAVHPA1 & NAVHPA 2 & NAVHPA3 & NAVHPA4 & NAVHPA5 & NAVHPA6 & NAVHPA7 (Solid State Power Amplifier)

ENVREQ-489 : NAVSW1 & NAVSW2 & NAVSW3 (RF High Pow er Switch):

Unit: NAVSW1 & NAVSW2 & NAVSW3 (RF High Power Switch)

T min Tmax


TCS Design -40 85


TCS Design -55 100






Table 62 - Temperature limits of NAVSW1 & NAVSW2 & NAVSW3 (RF High Power Switch)




ENVREQ-492 : NAVLOAD1 & NAVLOAD2 & NAVLOAD3 (RF Hi gh Power Load):

Unit: NAVLOAD1 & NAVLOAD2 & NAVLOAD3 (RF High Power Load)

T min Tmax


TCS Design -5 85


TCS Design -35 130






Table 63 - Temperature limits of NAVLOAD1 & NAVLO AD2 & NAVLOAD3 (RF High Power Load)

ENVREQ-495 : OPF1 &OPF2 (Output Filters):

Unit: OPF1 &OPF2 (Output Filters) T min Tmax


TCS Design -5 50


TCS Design -35 70





Table 64 - Temperature limits of OPF1 & OPF2 (Ou tput Filters)




ENVREQ-498 : OMUX (OMUX Diplexer LB (E5/E6)):

Unit: OMUX (OMUX Diplexer LB (E5/E6)): T min Tmax


TCS Design -5 50


TCS Design -35 70





Table 65 - Temperature limits of OMUX (OMUX Diple xer LB (E5/E6))

ENVREQ-501 : OPTC1 & OPTC2 & OPTC3 (Output Test Co uplers):

Unit: OPTC1 & OPTC2 & OPTC3 (Output Test Couplers)

T min Tmax


TCS Design -40 85


TCS Design -55 100






Table 66 - Temperature limits of OPTC1 & OPTC2 & OPTC3 (Output Test Couplers)




ENVREQ-504 : NAVANT (L-Band Navigation Antenna):

Unit: NAVANT (L-Band Navigation Antenna) T min Tmax


TCS Design -109 106


TCS Design -163 119

Start-up temperature (°C) N.A. None Ground storage and transport temp. (°C) -20 60 Stability requirement /short term None


Temperature limit definition (applicability) Main structure

Table 67 - Temperature limits of NAVANT (L-Band N avigation Antenna)

ENVREQ-507 : MISREC (C-Band Receiver):

Unit: MISREC (C-Band Receiver)

T min Tmax


TCS Design -10 45


TCS Design -35 70



T rel. (°C) 30




Table 68 - Temperature limits of MISREC (C-Band Receiver)




ENVREQ-516 : MISANT (C-band Mission Antenna):

Unit: MISANT (C-band Mission Antenna) T min Tmax


TCS Design -109 106


TCS Design -135 119

Start-up temperature (°C) N.A. None Ground storage and transport temp. (°C) -20 60



Temperature limit definition (applicability) Main Structure

Table 69 - Temperature limits of MISANT (C-band M ission Antenna)

ENVREQ-519 : SARANT (SAR Rx/Tx Antenna):

Unit: SARANT (SAR Rx/Tx Antenna) T min Tmax


TCS Design -109 106


TCS Design -135 119




Temperature limit definition (applicability) Main Structure

Table 70 - Temperature limits of SARANT (SAR Rx/T x Antenna)




ENVREQ-522 : SARIPTC (SAR Test Coupler 406 MHz):

Unit: SARIPTC (SAR Test Coupler 406 MHz) T min Tmax


TCS Design -40 85


TCS Design -55 100





Table 71 - Temperature limits of SARIPTC (SAR Tes t Coupler 406 MHz)

ENVREQ-525 : SAROPTC (SAR Test Coupler L-Band):

Unit: SAROPTC (SAR Test Coupler L-Band) T min Tmax


TCS Design -40 85


TCS Design -55 100





Table 72 - Temperature limits of SAROPTC (SAR Tes t Coupler L-Band)




ENVREQ-528 : SART (SAR Transponder Assembly):

Unit: SART (SAR Transponder Assembly) T min Tmax


TCS Design -10 40


TCS Design -35 80


T rel. (°C) 30




Table 73 - Temperature limits of SART (SAR Transp onder Assembly)

ENVREQ-531 : RTU (Remote Terminal Unit):

Unit: RTU (Remote Terminal Unit) T min Tmax


TCS Design -5 45


TCS Design -35 80


T rel. (°C) 30




Table 74 - Temperature limits of RTU (Remote Term inal Unit)




ENVREQ-534 : HPHAR01 till HPHAR017 (RF Cables)

Unit: HPHAR01 till HPHAR017 (RF Cables) T min Tmax


TCS Design -15 45


TCS Design -45 140





Table 75 - Temperature limits of HPHAR01 till HPH AR017 (RF Cables)

ENVREQ-537 : LPHAR (RF LP Harness):

Unit: LPHAR (RF LP Harness) T min Tmax


TCS Design -15 45


TCS Design -35 80





Table 76 - Temperature limits of LPHAR (RF LP Har ness)




ENVREQ-540 : CBHAR (C-Band Harness) and SARHAR (SA R Harness):

Unit: CBHAR (C-Band Harness) and SARHAR (SAR Harness)

T min Tmax


TCS Design -15 45


TCS Design -35 80






Table 77 - Temperature limits of CBHAR (C-Band Ha rness) and SARHAR (SAR Harness)

4.2.15.3 Temperature limits of S/C Structure Units

ENVREQ-544 : STPF-x, STPF+y, STPF-y, STPF-z, STPFM , STPFSF1, STPFSF2, STPL+x, STPL+y, STPL-y, STPL+z, S/C Structure Panel s:

STPF-x, STPF+y, STPF-y, STPF-z, STPFM, STPFSF1, STPFSF2, STPL+x, STPL+y, STPL-y, STPL+z, S/C Structure Panels:

T min Tmax


TCS Design -85 80


TCS Design -85 80

Start-up temperature (°C) NA NA




Temperature limit definition (applicability) Total Panel Structure

Table 78 - Temperature limits of STPF-x, STPF+y, STPF-y, STPF-z, STPFM, STPFSF1, STPFSF2, STPL+x, STPL+y, STPL-y, STPL+z, S/C Struct ure Panels




ENVREQ-547 : LSSIF1, LSSIF2, LSSIF3, LSSIF4, Separ ation Interface:

T min Tmax

Qualification -80 (TBC) 90 (TBC) Acceptance -75 (TBC) 85 (TBC)

Operating temperature (°C) (at release)

TCS Design -70 (TBC) 80 (TBC)

Qualification -110 (TBC) 110 (TBC) Acceptance -105 (TBC) 105 (TBC)

Non-Operating temperature (°C) (before and after release)

TCS Design -100 (TBC) 100 (TBC)


Ground storage and transport temp. (°C) -20 (TBC) 60 (TBC)



Temperature limit definition (applicability) Separation Plate on S/C

Table 79 - Temperature limits of LSSIF1, LSSIF2, LSSIF3, LSSIF4, Separation Interface

4.2.15.4 Temperature limits of TCS Units

ENVREQ-551 : External MLI:

External MLI: T min Tmax


TCS Design -190 130


TCS Design -190 130

Start-up temperature (°C) NA NA Ground storage and transport temp. (°C) -20 60



Temperature limit definition (applicability) MLI Surface (space facing)

Table 80 - Temperature limits of External MLI




ENVREQ-554 : External High Temperature MLI:

External High Temperature MLI: T min Tmax


TCS Design -190 240


Acceptance -195 245 Non-Operating temperature (°C)

TCS Design -190 240






Table 81 - Temperature limits of External High Te mperature MLI

ENVREQ-557 : Internal MLI:

Internal MLI: T min Tmax


TCS Design -190 130


TCS Design -190 130





Table 82 - Temperature limits of Internal MLI




ENVREQ-560 : OSR (Optical Surface Radiator):

OSR (Optical Surface Radiator): T min Tmax


TCS Design -150 110


TCS Design -145 110




Temperature limit definition (applicability) OSR Surface

Table 83 - Temperature limits of OSR (Optical Sur face Radiator)

ENVREQ-563 : Heaters (Foil Heaters):

Heaters (Foil Heaters): T min Tmax


TCS Design -105 140


TCS Design -105 140




Temperature limit definition (applicability) Heater Surface

Table 84 - Temperature limits of Heaters (Foil He aters)




ENVREQ-566 : Thermistors:

Thermistors T min Tmax


TCS Design -60 90


TCS Design -60 90





Table 85 - Temperature limits of Thermistors

ENVREQ-569 : Heatpipes:

Heat Pipes T min Tmax


Acceptance -55 70 Operating temperature (°C)

TCS Design -50 65


TCS Design -60 110 Start-up temperature (°C) N.A. None




Temperature limit definition (applicability) Heat Pipe Surface (TRP)

Table 86 - Temperature limits of Heatpipes




4.2.15.5 Temperature limits of Harness

ENVREQ-573 : PDHAR (Power and Data Handling Harnes s): Unit: PDHAR (Power and Data Handling Harness) T min Tmax


TCS Design -15 45


TCS Design -35 80





Table 87 - Temperature limits of PDHAR (Power and Data Handling Harness)

4.2.16 Spacecraft Thermal Interfaces for External Units

4.2.16.1 Interface Temperatures

ENVREQ-578 : Interface temperatures for external u nits The following interface temperatures are defined for the spacecraft structure to be used for external units.

Spacecraft Interface Temperatures for External Unit s in [°C]

Transfer-orbit Normal-orbit Safe mode Interface

cold cases hot cases cold cases hot cases cold cases hot cases

Panel (wall) -30 40 -20 40 -20 40 +X MLI external foil -150 145 -150 145 -150 145 Panel (wall) -30 40 -20 60 -20 50

-X MLI external foil -150 150 -150 145 -150 145 Panel (wall) -35 40 -20 60 -30 50 +/-

Y MLI external foil -150 145 -150 145 -150 145 Panel (wall) -30 40 -20 60 -20 50

+Z MLI external foil -150 145 -150 145 -150 145 Panel (wall) -35 40 -20 60 -20 50

-Z MLI external foil -150 145 -150 145 -150 145

Note: all values are TBC

Table 88 - Interface temperatures for external un its




4.2.16.2 Thermo Optical Properties of Spacecraft E xternal Surfaces

ENVREQ-582 : Thermo optical properties for externa l units The following thermo optical properties of the spacecraft external surfaces shall be used for thermal analysis of external units.

Spacecraft Thermo Optical Properties (external) Absorptivity Emissivity

Location Coating BOL EOL

Panel (wall) OSR 0,08 0,25 0,8 +X MLI external foil MLI (black) 0,94 0,94 0,83

-X MLI external foil MLI (black) 0,94 0,94 0,83 Panel (wall) OSR 0,08 0,25 0,8

+/-Y MLI external foil MLI (black) 0,94 0,94 0,83

front side = cell side (GaAs) 0,92 0,92 0,82 +/-Y Solar Array

rear side = CFRP 0,9 0,9 0,72 +Z MLI external foil MLI (black) 0,94 0,94 0,83 -Z MLI external foil MLI (black) 0,94 0,94 0,83

Note: all values are TBC

Table 89 - Thermo optical properties for external units

ENVREQ-585 : Irradiative coupling of external MLI and satellite structure The effective irradiative coupling between external MLI foil and S/C structure walls shall be considered equal to: for large MLI blankets covering wide plane areas: b= 0,02 W/m²K and εeff = 0,011 (TBC) for small MLI blankets: covering small plane areas: b= 0,05 W/m²K and εeff = 0,011 (TBC)




5 Cleanliness

ENVREQ-587 : Cleanliness Conditions The following practices shall be followed so that the cleanliness conditions for the S/C are ensured: • Precautions shall be taken and a clean environment shall be provided during the

production, test and delivery of the S/C and all S/C related items • During all S/C ground activities organic deposits shall not exceed 4 mg/m²/week. • All spacecraft AIT activities shall be carried out in class 100 000 or better clean rooms. • Prior to the encapsulation of the S/C, the upper stages and fairing shall be cleaned and

their cleanness shall be checked. • All handling equipment shall be able to be introduced in clean-room and shall therefore

be cleaned and inspected before their entry in the facilities. • Once encapsulated, the upper composites shall be hermetically closed or a class 100

000 air-conditioning of the fairing shall be provided.

Filtration 0,3 µm

Noise level due to ventilation dB

max 95

Table 90 - Cleanliness Conditions




6 Electromagnetic Environment

ENVREQ-592 : Electromagnetic Environment The satellite and the satellite units shall be compliant with the EMC design and test requirements of [AD 01].

ENVREQ-593 : Magnetic on-ground environment During payload module level and satellite level ground operations (including payload module and satellite transportation as well as launch site operations), the magnetic field generated by the ground environment shall never exceed 3 Gauss at any distance lower or equal to 0.5 m from the satellite (or payload module) whatever the satellite (or payload module) state (on or off) is. This shall be verified my measurement preferably at any time and at least before any change of the satellite (or payload module) configuration (e.g. transfer to another test facility, transfer to container, etc.).

ENVREQ-594 : Applicability of above requirement Note: This requirement shall be taken into account in the design of the GSE, in the definition of the test set up (test facilities, distance between GSE and satellite, etc.) as well as in the design of the payload module and satellite containers used during transportation (including on the launch site).




7 Orbital Environment

ENVREQ-596 : Orbital Environment The satellite and the satellite units shall be compliant with the orbital environment specified in [AD 02].




8 Qualification and Acceptance Environmental Tests The objective of the qualification tests is to demonstrate the stability and integrity of the design when subjected to the rigor of the anticipated environmental stresses.

The qualification environmental tests shall simulate conditions or effects which are more severe than mission environments in order to provide assurance of detecting design deficiencies.

The objective of the acceptance test is to provide assurance of detecting manufacturing deficiencies. The acceptance environmental tests will simulate conditions of effects which are similar to the mission environment.

In this section the qualification and acceptance test requirements for the satellite level tests as well as for the unit level tests are defined.

The test requirements described herein shall be used as the basis for the preparation of detailed test procedures by the party, responsible for the test. Following each test, the unit shall be examined and operated for evidence of no damage and such evidence shall be documented.

The environmental tests and applicability on a specific test item are specified in the relevant specifications and/ or in the AIT plan and described in the relevant test procedures.

8.1 Unit Level Tests This section presents the environmental test requirements which are applicable at unit level.

8.1.1 Definition and Objective Unit level is the lowest hardware level subjected to environmental qualification and acceptance tests. All testing performed on levels lower than unit level (e.g. PCB, component etc.) is considered as development testing.

8.1.1.1 Qualification Tests Units can be qualified in one of the three following ways:

ENVREQ-601 : Qualification testing At least one unit of a set of identical flight standard units must satisfactorily survive a set of tests at least as severe as the qualification tests specified in the following sections.

ENVREQ-602 : Proto-flight qualification testing An engineering model shall be subjected and shall survive the qualification tests specified in the following sections. Thereafter, one unit of a set of identical flight standard units must satisfactorily survive the proto-flight qualification tests specified in the following sections.




ENVREQ-603 : Similarity If the unit is identical to that used a previous programme, and if an identical unit was qualification tested by this programme, to a set of tests at least as severe as the qualification tests specified in the following sections, qualification by similarity can be accepted, pending the space segment approval at EQSR.

8.1.1.2 Acceptance Tests The objective of the acceptance tests is to verify the ability of the flight units to be submitted to the flight environment and to verify the unit workmanship.

8.1.2 Test Sequence

ENVREQ-606 : Test sequence applicable for units The following table presents the environmental tests sequence for all units and makes reference to the sections where the requirements are defined. Suppliers wishing to modify the order of the sequence for qualification, proto-qualification and acceptance testing shall ask for prior approval to the space segment prime. For units which have a sensitivity to thermal cycles (mechanical degradation), thermal test may have to be performed before mechanical test (agreement between unit supplier and space segment prime).

QUALIFICATION SEQUENCE

PROTOFLIGHT SEQUENCE

ACCEPTANCE SEQUENCE

SECTION

Full performance check Sinus vibration Good health check Random or acoustic Good health check Constant acceleration (5) Limited performance check Depressurisation Corona & Arcing Thermal Vacuum Mechanical measurement Shock test (2) EMC ESD Full performance check Life test

X X X X X X X X X X X X X X X X

X X X X X X X X X X X

X (4)

X

X

X

X

If applicable X X

(1)

X

AD 01 AD 01

(1) Limited EMC test as required in the unit specification. In particular each RF flight unit must be EMC tested (radiated emission and susceptibility), to detect any workmanship defect. All units shall be tested w.r.t. conducted emissions including measurements of the inrush current. (2) Shock test may be performed right after constant acceleration test pending Prime Contractor approval. (3) deleted (4) According to the test results on the QM, additional tests might be performed. (5) May be grouped with sine vibration as explained in §10.2.

Table 91 - Nominal Unit Test Sequence




ENVREQ-609 : All qualification testing shall be completed prior to the first flight. Environmental acceptance testing shall be completed prior to installation on a flight spacecraft and prior to qualification/ proto-qualification testing at spacecraft level.

8.1.3 Deleted

8.1.4 General test conditions

8.1.4.1 Ambient Conditions

ENVREQ-613 : Ambient test conditions Unless otherwise specified, all measurements and tests performed at ambient shall be made within the following conditions:

• -temperature : 23° + 5°C • -relative humidity : 25% - 60 % • -pressure : 760 + 25 mm mercury

Whenever these conditions must be closely controlled in order to obtain reproducible results, a reference temperature of 23 °C, a relative humidi ty of 50 % and an atmospheric pressure of 760 mm of mercury respectively shall be used, together with whatever tolerances are required to obtain the desired precision of measurement.

Actual ambient test conditions shall be recorded periodically during the test period.

Should the limits on ambient conditions of temperature and pressure be exceeded, the decision not to test or to stop any test in progress shall be vested in the test conductor, who must ensure that there will be no adverse influences on unit performance. However, the temperature of the unit shall not be allowed to exceed the relevant range. Decisions made by the test conductor must subsequently be ratified by the next Non conformance Review Board (NRB).

8.1.4.2 Tolerance of Test Conditions

ENVREQ-615 : The maximum allowable tolerances of all test conditions except those tests performed at ambient and specified in the previous section (exclusive of accuracy of instruments), shall be as specified in RD 09.

8.1.4.3 Measurements

ENVREQ-617 : The accuracy of all instruments and test equipments used to control or monitor the test parameters shall be verified periodically. All instruments and test equipments used to conduct the specified tests shall have an accuracy of better than one third of the tolerance for the variable to be measured.




8.1.4.4 Rejection and Retest

ENVREQ-619 : Rejection and Retest If a failure, malfunction or out-of-tolerance performance occurs during or after a test, the test shall be stopped and, the deficiency addressed in a dedicated Non conformance Review Board (NRB), as part of the general non conformance system. A complete record shall be made of any such failures and the corrective actions taken.

8.1.4.5 Change on a Unit

ENVREQ-621 : Change on a Unit Any change on a unit under a formal test sequence (qualification or acceptance) shall be submitted to a dedicated Non conformance Review Board (NRB) and formally approved by the space segment prime.

8.1.5 Performance checks between and during tests

ENVREQ-623 : Performance checks Performance checks and inspections shall be performed alternately or simultaneously with the environment tests. Three levels of performance check are defined: • "Full Performance Check" (FPC), which includes measurement of all the parameters

which are necessary to demonstrate compliance of the unit with required performance, • "Limited Performance Check" (LPC), which includes a slightly reduced number of tests, • "Good Health Check" (GHC) is a rapid check of some critical parameters which is

performed to demonstrate with a good probability that no failure occurred and that the unit status was not modified during the environment test.

ENVREQ-624 : Measurements The list of measurements to be performed during each of these performance checks is provided in the relevant unit specification. Performance checks and environmental tests shall be alternately performed as specified in ENVREQ-606. The only exceptions are those items which cannot be realistically tested in ambient conditions.

ENVREQ-626 : Exceptions In such cases, initial testing shall be designed to prove compliance as far as possible without causing damage to the tested item. The exceptions allowed are all the items associated with fluid propellants, for which a substitute may be used to ensure operation (open/close) of the unit. For thruster performance, operation using the correct propellant in vacuum is necessary. In addition to the performance checks during the test which are specified in the relevant environmental test section, a GHC and a visual inspection shall be performed after each installation on the test facility and prior to removal of the item from the test facility.




8.1.6 Mechanical Tests

ENVREQ-628 : Test condition for units The tested unit shall be installed in the test facility at room temperature in a manner which is representative of the flight conditions. Plugs, covers, harness, tubing, etc., used in flight shall be in place. When mechanical or electrical connections are not used, the connections normally protected in flight shall be adequately covered.

ENVREQ-629 : Attachment to the test fixture Units shall be attached to the test fixture using flight representative bolts and the bolts shall be tightened to the same torque as foreseen for the flight.

ENVREQ-630 : The interface thermal filler, if any, is not requested for mechanical tests.

ENVREQ-631 : Grounding during test The grounding of the unit under test and of the circuit interfaces in the test bench shall be submitted to the space segment prime for approval, in order to check the representativity with respect to the unit grounding within the satellite.

ENVREQ-632 : Energized units during vibration test For units which are ON during launch: at all times during any vibration, acoustic and constant acceleration tests, the unit under test shall be energised and placed in the same functional configuration as intended for the launch. Power supply voltages, currents and selected performance parameters shall be continuously monitored and recorded with adequate bandwidth.

ENVREQ-633 : Energized units during shock tests For units which are ON during shock events (separation or solar array shocks): at all times during shock tests, the unit under test shall be energised and placed in the same functional configuration as intended for the shock events. Power supply voltages, currents and selected performance parameters shall be monitored during the test and it shall be verified no abnormal transient occurs during the test. Furthermore, it shall be verified that the functional configuration of the unit remains identical after the application of the shock.

ENVREQ-634 : Vibration survey A vibration survey (or resonance search) 0.5 g, 10 to 2000 Hz shall be conducted prior and after performing each various mechanical tests specified. For the best correlation of model, the low level signature can be done without harness.




ENVREQ-635 : Validation of connector assemblies For units which are not mechanically tested with a flight configuration with respect to the connector and harness configuration, the connector assemblies will be validated mechanically by a dedicated test, by similarity, or by local test.

ENVREQ-636 : Pressurized units Pressurised units shall be pressurised with the maximum expected operating pressure (MEOP) according to the launch condition.

ENVREQ-637 : Mechanical Test Report The unit supplier shall provide the mechanical test report. The purpose of this document is to demonstrate the compliance of the unit w.r.t. the requirements of the present document. It also aims at insuring a good correlation between the analysis and test results. As a guide-line, the table of contents of this document may be: 1. Introduction 2. Applicable and Reference Documents (test specification) 3. Description: • Model description (unit model, mass, reference, axes, etc.) • Test facilities (short description, date of the tests, etc.) • Instrumentation plan (attached in annex, with the axis of each sensors) 4. Test Sequence: included aborts, actual schedule, problems, etc. 5. Test Results Frequencies, Damping factors, • Finite Element Models comparison and explanations of the differences, if any • Quasi-static ; Sine, Random & Shock tests : o levels measured for the unit, axis per axis ; o comparison between tested and expected/qualified levels for each parts ; o comparison between low levels before/ after qualification test ; o final visual or other inspections. 6. Electrical and/or Alignment and/or dismounting « Good Health Checks » 7. Points of attention (if any) 8. Conclusion 9. Recommendations for Spacecraft test:

• notching criterion and test instrumentation sensors, if any ; • test instrumentation sensors to be checked, etc.

ANNEX: • picture of the unit on the test fixture ; qualification levels at the basis for each axis ; • important sensor curves, etc.

8.1.6.1 General

ENVREQ-663 : Testing of units with support bracket s The following levels apply at the unit interface to the satellite structure. For units which include a support bracket in the responsibility of the unit supplier, they apply at the structure to bracket interface (bracket considered as part of the unit). In this particular case, the unit supplier shall qualify the sub-assembly bracket plus unit to the levels specified hereafter.




The unit supplier may qualify the bracket and the unit separately but in this case the unit supplier shall provide the adequate justification which shall be submitted to the space segment prime for approval.

ENVREQ-664 : Test conditions for mechanical tests The following conditions shall apply to the mechanical tests: The input levels defined in this section are applicable at the test fixture/unit interface. To avoid significant amplification of inputs to the units during sine and random vibration test, the test fixture shall be designed taking account the following rules: • The first frequency of the test fixture loaded by the unit (or by equivalent masses and

inertia) shall be at least 1.5 times the first frequency of the unit alone. • The main frequencies of the test fixture loaded by the unit (or by an equivalent mass)

shall not be coupled with the main unit frequencies.

ENVREQ-665 : Success criteria Successful completion of the vibration survey shall be a requirement for qualification. The transmissibility frequency behaviour, and cross-talk of the test fixture loaded by the unit shall be determined by a swept sine or a low level random vibration applied in each of the three mutually perpendicular axes. The shift in frequency, if any, between the 2 low level tests before/after each mechanical test should be less than 5%. The variation of transmissibility between test item mounting points shall not exceed the factor of ± 3dB between 5 and 500 Hz and ± 6 dB between 500 and 2000 Hz, provided that the total cumulative bandwidth which exceeds ± 3 dB does not exceed 300 Hz. Cross talk shall not exceed the input. These values shall be measured with a sufficient number of accelerometers mounted at the unit/test fixture interface. For units sensitive to the in-orbit micro-vibrations levels specified in section 4.1.2.3.2, a low level run (as low as possible) shall be performed from 10 to 2000 Hz, the unit being operating and the sensitive parameters recorded. The levels to be used for demonstration of compliance are 1.5 times those of section 4.1.2.3.2 for qualification and equal to those of section 4.1.2.3.2 for acceptance.

ENVREQ-666 : General test level definition The quasi-static, sine and random levels are defined in the following sections according to the following table which defines the applicable zones for each satellite unit.

Unit Description Location (*) Quasi-static zone Sine zone Random zone Shock zone TTC TX/RX S-band transponder PF-Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-727, 730 ENVREQ-746 TTCHYB Hybrid coupler PF-Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-727, 730 ENVREQ-746 TTCANT TTC antenna PL+Z/PL+Y (E) ENVREQ-672 c) ENVREQ-682 ENVREQ-734 ENVREQ-756 TTCHAR TTC harness All zones (I, E) ENVREQ-672 a) ENVREQ-676 ENVREQ-706, 709 ENVREQ-744 LRR Laser Retro-reflector PL+Z (E) ENVREQ-672 a) ENVREQ-676 ENVREQ-720, 723 ENVREQ-750 PFSU Platform Security Unit PF-Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-727, 730 ENVREQ-746 EPS PCDU Power Conditioning Unit PF+Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-727, 730 ENVREQ-746 SAW Solar Array Wing PL+Y/PL-Y (E) Specific Specific n.a. (cf. acoustic) ENVREQ-746 SADM Solar Array Drive Mechanism PL+Y/PL-Y (E) ENVREQ-672 a) ENVREQ-676 ENVREQ-713, 716 ENVREQ-746 SAHD Solar Array Hold Down PL+Y/PL-Y (E) Specific Specific n.a. (cf. acoustic) ENVREQ-746 BATT Battery PF+Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-727, 730 ENVREQ-746




Avionics ICDU Integrated Control and Data handling Unit PF-Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-727, 730 ENVREQ-746 FSS Fine Sun Sensor PL+Z (E) ENVREQ-672 b) ENVREQ-679 ENVREQ-734 ENVREQ-756 CSS Coarse Sun Sensor PL+Z/PL-Y (E) ENVREQ-672 b) ENVREQ-679 ENVREQ-734 ENVREQ-756 ES Earth Sensor PL+Z (E) ENVREQ-672 a) ENVREQ-676 ENVREQ-720, 723 ENVREQ-750 GYRO Gyro PF-Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-734 ENVREQ-746 MTR Magneto-torquer Rod PF-X/PF-M (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-706, 709 ENVREQ-744 RW Reaction Wheels PF-M (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-974, 977 ENVREQ-754 Propulsion PROTANK Tank PF-SF1/2 (I) ENVREQ-672 d) ENVREQ-685 ENVREQ-727, 730 ENVREQ-754 PROFVV Fill and Vent Valve PF-X (E) ENVREQ-672 a) ENVREQ-676 ENVREQ-706, 709 ENVREQ-744 PROPT Pressure Transducer PF-X (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-706, 709 ENVREQ-744 PROLF Filter PF-X (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-706, 709 ENVREQ-744 PROFDV Fill and Drain Valve PF-X (E) ENVREQ-672 a) ENVREQ-676 ENVREQ-706, 709 ENVREQ-744 PROLV Latch Valve PF-X (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-706, 709 ENVREQ-744 PROTP Test Port PF-X (E) ENVREQ-672 a) ENVREQ-676 ENVREQ-706, 709 ENVREQ-744 PRORCT Reaction Control Thruster PF-X (E) ENVREQ-672 a) ENVREQ-676 ENVREQ-706, 709 ENVREQ-744 Harness Harness Harness hardware All zones (I, E) ENVREQ-672 a) ENVREQ-676 ENVREQ-706, 709 ENVREQ-744 Thermal Thermal Thermal hardware All zones (I,E) ENVREQ-672 a) ENVREQ-676 ENVREQ-706, 709 ENVREQ-744 Payload RAFS Rubidium Atomic Frequency Standard PL+X (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-699, 702 ENVREQ-742 PHM Passive Hydrogen Maser PL+X (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-967, 970 ENVREQ-742 CMCU Clock Monitoring and Control Unit PL-Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-713, 716 ENVREQ-746 PLSU Payload Security Unit PL+Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-713, 716 ENVREQ-746 NSGU Navigation and Signal Generator Unit PL+Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-713, 716 ENVREQ-748 FGUU Frequency Generation and Up-conversion Unit PL+Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-713, 716 ENVREQ-746 FRTC Frequency Reference Test Coupler PL+Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-713, 716 ENVREQ-746 CRTC Clock Reference Test Coupler PL+Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-713, 716 ENVREQ-746 BBTC Base Band Test Coupler PL+Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-713, 716 ENVREQ-746 IPTC Input Test Coupler PL+Z (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-720, 723 ENVREQ-750 SPLIT Splitter PL+Y/PL+Z (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-720, 723 ENVREQ-746 NAVHPA Navigation SSPA PL+Y/PL-Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-713, 716 ENVREQ-748 NAVSW High Power Switch PL+Z (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-720, 723 ENVREQ-750 NAVLOAD High Power Load PL+Y/PL-Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-713, 716 ENVREQ-748 OPF Output Filter PL+Z (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-720, 723 ENVREQ-750 OMUX Output Multiplexer PL+Z (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-720, 723 ENVREQ-750 OPTC Output Test Coupler PL+Z (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-720, 723 ENVREQ-750 NAVANT Navigation Antenna PL+Z (E) Specific specific n.a. (cf. acoustic) ENVREQ-750 MISREC Mission receiver PL-Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-713, 716 ENVREQ-746 MISSP C-band splitter PL+Z (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-720, 723 ENVREQ-750 MISTC C-band test coupler PL+Z (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-720, 723 ENVREQ-750 MISANT C-band antenna PL+Z (E) ENVREQ-672 a) ENVREQ-676 ENVREQ-720, 723 ENVREQ-750 SARANT SAR antenna PL+Z (E) Specific specific n.a. (cf. acoustic) ENVREQ-750 SARIPTC SAR Input Test Coupler PL+Z (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-720, 723 ENVREQ-750 SAROPTC SAR Output Test Coupler PL+Z (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-720, 723 ENVREQ-750 SART SAR Transponder PL+Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-713, 716 ENVREQ-748 RTU Remote Terminal Unit PL-Y (I) ENVREQ-672 a) ENVREQ-676 ENVREQ-713, 716 ENVREQ-746 HPHAR High Power Harness All zones (I, E) ENVREQ-672 a) ENVREQ-676 ENVREQ-706, 709 ENVREQ-744 LPHAR Low Power Harness All zones (I, E) ENVREQ-672 a) ENVREQ-676 ENVREQ-706, 709 ENVREQ-744 CBHAR C-band Harness All zones (I, E) ENVREQ-672 a) ENVREQ-676 ENVREQ-706, 709 ENVREQ-744 SARHAR SAR harness All zones (I, E) ENVREQ-672 a) ENVREQ-676 ENVREQ-706, 709 ENVREQ-744

(*) (I): Internally accommodated unit, (E) externally accommodated unit.

Table 92 - Quasi-static, sine and random levels

8.1.6.2 Quasi-static test

ENVREQ-670 : Quasi-static test All units shall be designed to withstand the following specified quasi-static loads. This shall be verified by test.




Qualification and proto-qualification levels are identical and the test duration is 60 s in both cases for each axis. Qualification w.r.t quasi-static loads may be performed by a sine vibration test performed at a frequency well below the unit first resonance frequencies. In this case, a constant frequency sine input (typical value: 35 Hz) shall be held for one minute after the steady state level has been established. In the following the applicable quasi-static levels are provided. Specific levels have been defined for units which are mounted on brackets at satellite level.

ENVREQ-672 : Qualification quasi-static loads The following qualification quasi-static loads shall apply: a) For all units except those mentioned hereafter: 20 g b) Sun Sensors: 30 g c) TTC antenna: 50 g d) Propellant tank: 16 g along mounting axis / 12 g lateral For large appendages (solar array, navigation antenna, SAR antenna), dedicated quasi-static loads will be provided by the space segment prime. These loads will be confirmed by the satellite dynamic analyses which will be performed using the reduced dynamic mechanical models delivered by the corresponding unit suppliers.

8.1.6.3 Sine Test The objective of this test is to verify that the minimum design stiffness requirements are met and to verify that the design can survive later spacecraft level sine vibration testing.

ENVREQ-675 : Sine levels The following levels shall be applied separately along the three orthogonal axes. Qualification levels and proto-qualification levels are identical with the following test sweep rates (starting from lower frequency): \'b7 2 octaves/minute for qualification test \'b7 4 octaves/minute for proto-qualification test

ENVREQ-676 : Sine qualification levels For all units except those specified hereafter, the following qualification levels shall apply (all three axes):

FREQUENCY (Hz) ACCELERATION (g 0-peak) 5 - 20

20 - 100 10 mm 0-peak (or max. shaker limit)

20 g

Table 93 - Sine qualification levels

ENVREQ-679 : Sun sensor For the sun sensors, the following qualification levels shall apply (all three axes):




FREQUENCY (Hz) ACCELERATION (g 0-peak)

5 - 20 20 - 100

10 mm 0-peak (or max. shaker limit) 30 g

Table 94 - Sun Sensors sine qualification levels

ENVREQ-682 : TTC antennas For the TTC antennas, the following qualification levels shall apply (all three axes):

FREQUENCY (Hz) ACCELERATION (g 0-peak) 5 - 20

20 - 100 10 mm 0-peak (or max. shaker limit)

50 g

Table 95 - TTC antennas sine qualification levels

ENVREQ-685 : Propellant tank For the propellant tank, the following qualification levels shall apply: Longitudinal (along the mounting axis):

FREQUENCY (Hz) ACCELERATION (g 0-peak) 5 – 22.6

22.6 - 100 6.4 mm 0-peak (or max. shaker limit)

16 g

Table 96 - Propellant tank sine qualification lev els (longitudinal)

Lateral (normal to the mounting axis): FREQUENCY (Hz) ACCELERATION (g 0-peak)

5 – 18.8 18.8 - 100

6.4 mm 0-peak (or max. shaker limit) 12 g

Table 97 - Propellant tank sine qualification lev els (lateral)

ENVREQ-692 : Large appendages For the large appendages (solar array, navigation antenna, SAR antenna), dedicated sine levels will be provided by the space segment prime. These levels will be confirmed by the satellite dynamic analyses which will be performed using the reduced dynamic mechanical models delivered by the corresponding unit suppliers. For these units which eigen-frequencies may be below 100 Hz, a notching procedure may be used. This shall however be considered as a waiver which shall be submitted to the space segment prime for approval.




8.1.6.4 Random Test

ENVREQ-694 : Random loads The random loads as defined below shall be taken into account to demonstrate the ability of the units to withstand the random excitation produced by the launcher and the transmitted acoustic noise excitations.

ENVREQ-695 : Random Test Duration Qualification levels and proto-qualification (PQ) levels are identical. Test durations are: - 180 s for qualification test; - 60 s for proto-qualification; - 60 s for acceptance tests.

ENVREQ-696 : Random vibration levels The unit shall be submitted to the following random vibration input separately along each orthogonal axis. For units mounted directly on the satellite panels, the following random levels apply: For units located on the +X panel 1) RAFS

ENVREQ-699 : Out of plane:

QUALIF/PQ ACCEPTANCE 20 Hz – 100 Hz

100 Hz – 400 Hz 400 Hz – 2000 Hz

+3 dB / octave 0.5 g2 / Hz

-3 dB / octave


-3 dB / octave Overall level 22.27 gRMS 15.75 gRMS

Table 98 - Out of plane random vibration levels f or units located on the +X panel 1) RAFS

ENVREQ-702 : In plane:


100 Hz – 1000 Hz 1000 Hz – 2000 Hz


-3 dB / octave



Table 99 - In plane random vibration levels for u nits located on the +X panel 1) RAFS




For units located on the +X panel 2) PHM



100 Hz – 400 Hz 400 Hz – 2000 Hz


-5 dB / octave



Table 100 - Out of plane random vibration levels for units located on the +X panels 2) PHM



100 Hz – 600 Hz 600 Hz – 2000 Hz


-5 dB / octave



Table 101 - In plane random vibration levels for units located on the +X panels 2) PHM

For units located on the -X panels (propulsion units except tank):



100 Hz – 400 Hz 400 Hz – 2000 Hz


-6 dB / octave



Table 102 - Out of plane random vibration levels for units located on the -X panels (propulsion units except tank)






100 Hz – 500 Hz 500 Hz – 2000 Hz


-6 dB / octave



Table 103 - In plane random vibration levels for units located on the -X panels (propulsion units except tank)

For units located on the +/- Y payload panels:



100 Hz – 250 Hz 250 Hz – 2000 Hz


-6 dB / octave



Table 104 - Out of plane random vibration levels for units located on the +/- Y payload panels



100 Hz – 600 Hz 600 Hz – 2000 Hz


-6 dB / octave



Table 105 - In plane random vibration levels for units located on the +/- Y payload panels




For units located on the +Z panel:



100 Hz – 300 Hz 300 Hz – 2000 Hz


-6 dB / octave



Table 106 - Out of plane random vibration levels for units located on the +Z panel



100 Hz – 600 Hz 600 Hz – 2000 Hz


-6 dB / octave



Table 107 - In plane random vibration levels for units located on the +Z panel

For the units located on the platform +Y/-Y/ panels and shear frames (including tank):



100 Hz – 300 Hz 300 Hz – 2000 Hz


-6 dB / octave



Table 108 - Out of plane random vibration levels for units located on the platform +Y/-Y/ panels and shear frames






100 Hz – 600 Hz 600 Hz – 2000 Hz


-6 dB / octave



Table 109 - In plane random vibration levels for units located on the platform +Y/-Y/ panels and shear frames

For the units located on the platform M panel



100 Hz – 400 Hz 400 Hz – 2000 Hz


-3 dB / octave



Table 110 - Out of plane random vibration levels for units located on the platform M panel



50 Hz – 500 Hz 500 Hz – 2000 Hz


-4dB / octave



Table 111 - In plane random vibration levels for units located on the platform M panel

For the sun sensors and the TTC antennas, the following random levels apply:




ENVREQ-734 : All three axes:


100 Hz – 300 Hz 300 Hz – 2000 Hz


-6 dB / octave



Table 112 - All three axes random vibration level s for the sun sensors and the TTC antennas

8.1.6.5 Acoustic Test

ENVREQ-738 : Acoustic Noise Test Large units which are sensitive against acoustic excitation (e.g. solar arrays, navigation antenna, SAR antenna) shall be submitted to an acoustic noise test which shall be performed instead of the random test.

ENVREQ-739 : Acoustic test requirements • The following requirements apply: • The tested item shall be mounted on a flight type support structure with ground

handling equipment removed. • The tested item shall be in flight configuration. • The test chamber shall provide a uniform energy density. • The sound pressure flight test levels are defined in sections 4.1.1.2.4 and the test

factors for qualification and acceptance tests are defined in section 3.4, The test duration shall be:

• Qualification: 180 s • Proto-qualification: 60 s • Acceptance: 60 s • The sound pressure levels presented in this document represent the worst case

combination of the acoustic spectra of the candidate launchers.

8.1.6.6 Shock Test The spacecraft is submitted to shocks, principally during its separation from the launch vehicle upper stage and during the solar array hold-on point release. These shocks propagate within the spacecraft and generate at the unit interface the shock levels specified in this section.




ENVREQ-741 : Qualification shock levels The satellite units shall be designed to withstand without degradation the qualification shock levels defined in the following tables. The following shock response spectrums (SRS) are applicable to each axis independently.

ENVREQ-742 : For the units mounted on the +X panel (PFM, RAFS): All axes:

Frequency SRS 100 Hz 50 g 1000 Hz 2500 g

10000 Hz 2500 g

ENVREQ-744 : For the units mounted on the -X panel (propulsion units except tank): All axes:


10000 Hz 5000 g

ENVREQ-746 : For the units directly mounted on the +/-Y panels: All axes:


10000 Hz 5000 g

ENVREQ-748 : For the units mounted on heat-pipes o n the +/-Y panels: All axes:


10000 Hz 3600 g




ENVREQ-750 : For the units mounted on the +Z panel s: All axes:


10000 Hz 3600 g

ENVREQ-752 : For the units mounted on the -Z panel s: All axes:


10000 Hz 5000 g

ENVREQ-754 : For the reaction wheels and the tank: All axes:


10000 Hz 2500 g

ENVREQ-756 : For the sun sensors and TTC antennas: All axes:


10000 Hz 3600 g

No shock test is required in the unit level acceptance test programme.

8.1.7 Thermal Tests

8.1.7.1 Qualification Thermal Tests The following requirements apply to test articles subjected to thermal vacuum qualification testing. General definitions for the qualification temperature limits are given in chapter 4.2.1.5, for reference also see ECSS-E-10-03 [RD 07].




8.1.7.1.1 General requirements

ENVREQ-762 : Derivation of temperature limits The qualification temperature limits shall be derived from the results of the overall S/C thermal analysis and shall be detailed in the subsystem test requirement specifications and the unit test requirement specifications. In case of re-use of off-the-shelf unit, the qualification temperature limits of these units shall be respected by the thermal design as far as possible.

ENVREQ-763 : The unit supplier shall perform qualification tests with his unit under qualification operating temperature limits, in vacuum, with functioning, fulfilling all required performances and without performance degradation.

ENVREQ-764 : The unit supplier shall perform qualification tests with his unit under qualification non-operating temperature limits, in vacuum, without functioning and without performance degradation.

ENVREQ-945 : For internally redundant units, the unit supplier shall perform performance qualification test for the prime and redundant parts, with switching from one to the other.

ENVREQ-765 : The suitable value to account for test temperature inaccuracies or test temperature tolerance shall be added to the qualification temperature range. See chapter 4.2.1.5.

ENVREQ-766 : The unit supplier shall perform during his qualification tests with his unit the application of the start-up temperature without performance degradation.

ENVREQ-767 : The unit supplier shall perform qualification tests with his unit under ground storage temperature, without functioning and without performance degradation.

ENVREQ-768 : The temperatures shall be controlled measured and selected such, that it can be guaranteed that the test item sees the required qualification temperature range including the test tolerance.




ENVREQ-769 : Throughout the thermal testing, the performance shall be monitored to an extent necessary to detect any anomalous behaviour of the unit.

ENVREQ-770 : Within the thermal vacuum test during reduction of pressure, the presence of corona and arcing shall be determined for unit operating during launch. Corona and arcing shall not occur with the units operating at that time.

ENVREQ-771 : The test conductance shall be detailed in the subsystem test requirement specifications and the unit test requirement specifications.

ENVREQ-772 : The test set-up shall be detailed in the related test procedures for the test article.

8.1.7.1.2 Test arrangements for internally mounted 'radiative' units

ENVREQ-774 : Test arrangement for isothermal radia tive units The test arrangement required for the unit having temperature limits referred to as either: a) The average skin temperature (Tskin) as shown schematically in figure below or b) The average environment temperature (Tenv) as shown schematically in figure below. The unit shall be installed in a thermal vacuum chamber such that it is conductively isolated (i.e. either suspended or mounted on a conductively isolated support). The assembly shall radiate to a thermally controlled isothermal black painted uniform environment. The environment shall be provided by either the chamber shroud or an arrangement of thermal control panels. Sufficient instrumentation shall be provided to allow accurate temperature control of the thermal control panel(s) and/or shroud by using a fluid loop and/or electrical resistance heaters.

ENVREQ-775 : Control of the shroud During test, the shroud or other thermal control panels (radiative environment) shall be controlled to either: a) Give the required temperature level on the unit reference temperature (see unit specification) or b) Give directly the required temperature level on the radiative environment (see unit specification)




ENVREQ-776 : Test arrangement for non- isothermal radiative units For non-isothermal units the temperature of the environment (shroud) shall be monitored by the hottest point during hot phases and by the coldest point during cold phases.

ENVREQ-777 : Delta temperature between shroud and temperature reference point In any case (isothermal or non isothermal radiative unit), the maximum temperature gap between the shroud and the unit reference shall be less than 10°C for hot phases

Vacuum Chamber

Temperature controld Chamber Shroud (hight emittacne >0,8)

Conducutiveinsulated support

Equipment/Unit

Tskin

Tenv

Figure 23 - isothermal radiative units

8.1.7.1.3 Test arrangements for internally mounted 'conductive' units The test arrangement required for units having temperature limits referred to the average base-plate temperature (Tbase) is described hereby and shown in the figure below:

ENVREQ-781 : Attachment of the unit The unit shall be bolted directly on a thermally controlled conductive heat sink using correct bolt and bolt torques using a flight representative thermal interface filler (to be procured by the unit supplier).

ENVREQ-782 : Temperature control During the test the conductive heat sink temperature shall be controlled to give the required temperature level on the unit reference temperature (Tbase or previously agreed reference point, i.e. TRP).

ENVREQ-783 : Qualification temperature The qualification temperatures to be reached are defined in chapter 4.2.15

ENVREQ-784 : Control of the radiative heat exchang e During the test the radiative exchanges shall be thermally controlled with a MLI blanket placed on the unit (option A) or the chamber radiative shroud may be controlled to the same temperature as the unit base-plate (option B).




The choice between both options has to be submitted to space segment prime for approval. For non-isothermal units, the temperature of the heat sink shall be monitored by the hottest point during hot phases and by the coldest point during cold phases.

Vacuum Chamber

Conducutive InterfacePlate (Tbase)

Equipment/Unitconductive mounted

Tbase

OPTION A:

MLIVacuum Chamber

Temperature controld Chamber Shroud (hight emittacne >0,8)

Conducutive InterfacePlate (Tbase)

Equipment/Unitconductive mounted

Tenv = Tbase

OPTION B:

Figure 24 Test arrangements for internally mounted 'conductive' units

8.1.7.1.4 Test arrangement for special internal un its

ENVREQ-788 : Special test provisions Certain internally mounted unit will require special test provisions. Examples of such unit would be: • Sensors having viewing apertures seeing space. • High power unit mounted on heat pipes (SSPA for example) • Radiatively decoupled units In such cases the required approach is to modify the test methods given in the previous sections to the extent needed to give a reasonably representative test environment. Potential problems in this respect should be brought to the attention of the space segment prime in order to allow an agreement to be found prior to testing.

8.1.7.1.5 Test arrangement for externally mounted units

ENVREQ-790 : Design criteria for the test arrangem ent The test arrangement must be designed to give the required qualification or acceptance temperatures on the unit with approximately representative heat flows to and from the environment. Where heat flows to and from the external environment predominate, the equipment shall be tested under space simulation conditions with the test chamber cold wall at liquid nitrogen (LN2) temperature and an infra-red radiation source equivalent to at least one solar constant




in the test plane to control temperatures over the required ranges for at least 8 cycles for qualification and 4 cycles for acceptance.

8.1.7.1.6 Test sequence requirements

ENVREQ-792 : Vacuum requirement The unit shall be tested in a thermal vacuum environment having a pressure of 1E-5 hPa or less.

ENVREQ-793 : Qualification test cycles The number of cycles applicable to the TV qualification test shall be n = 8.

ENVREQ-794 : Start of the test The start cycle shall be a hot cycle to achieve out-gassing of the unit.

ENVREQ-795 : Number of performance tests The unit shall be performance tested at least 4 times during the thermal vacuum cycling. These steps are in full vacuum at: • Ambient on start of testing • Hot plateau • Cold plateau • Ambient on end of testing

ENVREQ-796 : Temperature over time gradient for sa tellite internal units For units inside the S/C the slope (temperature/time gradient) on the unit shall be dT/dt ≤ 2°C per minute

ENVREQ-797 : Temperature over time gradient for sa tellite external units For units outside the S/C the slope (temperature/time gradient) on the unit shall be dT/dt ≤ TBD °C/ per minute (TBD in the unit specification).

ENVREQ-798 : Temperature stability criteria The temperature stabilisation shall be defined as achieved if the average temperatures of thermo couples on the unit have reached the slope (temperature/time gradient) of 1°C/1 hour and the acceptance temperature limit at a range of 0°C to -3°C for the cold case or 0°C to +3 for the hot case. After stabilisation, the temperature is to be held constant for the dwell time.

ENVREQ-799 : Dwell time The minimum stabilised temperature time shall be 2 hours = dwell time.




ENVREQ-800 : Test sequence The following test sequence (see figure next page) shall apply: • Unit functional performance test at ambient (temperature and pressure) • Unit off, heating up to maximum non operating qualification temperature • Temperature stabilisation and dwell time for 2 hours • Unit cooling down to minimum non operating qualification temperature • Temperature stabilisation and dwell time for 2 hours • Heating up to start-up temperature and unit switch on • Verification of unit switching capability by switching off and on at start-up temperature • Unit heating up in on mode to maximum operating qualification temperature • Temperature stabilisation and dwell time for 2 hours • Verification of unit hot switching capability by switching off and on at maximum

operating qualification temperature • Unit cooling down in on mode to minimum operating qualification temperature • Temperature stabilisation and dwell time for 2 hours • Unit heating up and cooling down to perform the required number of cycles • On last thermal vacuum cycle: unit performance test on maximum operating

qualification temperature • On last thermal vacuum cycle: unit performance test on minimum operating

qualification temperature • Unit heating up to ambient temperature in vacuum • Unit switch off at ambient temperature in vacuum • Ambient pressure recovery Unit functional performance test at ambient (temperature and pressure) Furthermore, at all times where the unit is on, a selected set of parameters shall continuously be monitored (typically one measurement per minute).




T ambient

TNF max.

Nomenclatur:

T ambient ambient temperature

TNFmax. maximum non-functioning temperature

TNFmin. miniumm non-functioning temperature

TFQmax. maximum qualification temperature

TFQmin. minimum qualification temperature

TSUmin. minimum start-up temperature

P pressure of TV chamber

tS dwell time = minimum stabilised temperature time = 2hours

tE minimum stabilised temperature time prior to start performance test: 4 hours

1 performance test

2 unit on/off (verification of on/off switching capability at TSU and TFQ)

3 at all the times when unit is on, a selected set of major parameters shall continously

monitored (typically 1 point per minute)

4 for Corona test, some units are on during pressure drop

TFQmax.

TFQmin.

TNFmin.

TSUmin.

tS tS

tS tS

tS

tE

tE

P

Time

1

1

1

1

2

3

Unit on for corona test

Unit off

Unit on

2

tS tS tS tS

tS tS tS

tS

tS

4

Figure 25 Qualification Thermal Test Sequence

8.1.7.2 Acceptance Thermal Test The following requirements apply to test articles subjected to thermal vacuum acceptance testing. General definitions for the acceptance temperature limits are given in chapter 4.2.1.5, for reference also see ECSS-E-10-03 [RD 07].




8.1.7.2.1 General requirements

ENVREQ-805 : Derivation of the acceptance temperat ure limits The acceptance temperature limits shall be derived from the results of the overall S/C thermal analysis and shall be detailed in the subsystem test requirement specifications and the unit test requirement specifications. In case of re-use of off-the-shelf unit, the acceptance temperature limits of these units shall be respected by the thermal design as far as possible.

ENVREQ-806 : The unit supplier shall perform acceptance tests with his unit under acceptance operating temperature limits, in vacuum, with functioning, fulfilling all required performances and without performance degradation.

ENVREQ-807 : The unit supplier shall perform acceptance tests with his unit under acceptance non-operating temperature limits, in vacuum, without functioning and without performance degradation.

ENVREQ-946 : For internally redundant units, the unit supplier shall perform performance acceptance test for the prime and redundant parts, with switching from one to the other.

ENVREQ-808 : The suitable value to account for test temperature inaccuracies or test temperature tolerance shall be added to the acceptance temperature range. See chapter 4.2.1.5.

ENVREQ-809 : The unit supplier shall perform during his qualification tests with his unit the application of the start-up temperature without performance degradation.

ENVREQ-810 : The unit supplier shall perform acceptance tests with his unit under ground storage temperature, without functioning and without performance degradation.

ENVREQ-811 : The temperatures shall be controlled measured and selected such, that it can be guaranteed that the test item sees the required acceptance temperature range including the test tolerance.




ENVREQ-812 : Throughout the thermal testing, the performance shall be monitored to an extent necessary to detect any anomalous behaviour of the unit.

ENVREQ-813 : Within the thermal vacuum test during reduction of pressure, the presence of corona and arcing shall be determined for unit operating during launch. Corona and arcing shall not occur with the units operating at that time.

ENVREQ-814 : The test conductance shall be detailed in the subsystem test requirement specifications and the unit test requirement specifications.

ENVREQ-815 : The test set-up shall be detailed in the related test procedures for the test article.

ENVREQ-816 : The test arrangements for acceptance testing shall be the same as for qualification testing.

8.1.7.2.2 Test sequence requirements

ENVREQ-818 : The unit shall be tested in a thermal vacuum environment having a pressure of 1E-5 hPa or less.

ENVREQ-819 : The number of cycles applicable to the TV acceptance test shall be n= 4.

ENVREQ-820 : The unit shall be performance tested at least 4 times during the thermal vacuum cycling. These steps are in full vacuum at: • Ambient on start of testing • Hot plateau • Cold plateau • Ambient on end of testing •

ENVREQ-821 : For units inside the S/C the slope (temperature/time gradient) on the unit shall be dT/dt ≤ 2°C per minute




ENVREQ-822 : For units outside the S/C the slope (temperature/time gradient) on the unit shall be dT/dt ≤ TBD °C/ per minute (TBD in the unit specification).

ENVREQ-823 : The minimum stabilised temperature time shall be 1 hour = dwell time.

ENVREQ-824 : The temperature stabilisation shall be defined as achieved if the average temperatures of thermo couples on the unit have reached the slope (temperature/time gradient) of 1°C/1 hour and the acceptance temperature limit at a range of 0°C to -3°C for the cold case or 0°C to +3 for the hot case. After stabilisation, the temperature is to be held constant for the dwell time.

ENVREQ-825 : The following test sequence (see figure next page) shall apply: • Unit functional performance test at ambient (temperature and pressure) • Unit off, heating up to maximum non operating acceptance temperature • Temperature stabilisation and dwell time for 1 hours • Unit cooling down to minimum non operating acceptance temperature • Temperature stabilisation and dwell time for 1 hours • Heating up to start-up temperature and unit switch on • Verification of unit switching capability by switching off and on at start-up temperature • Unit heating up in on mode to maximum operating acceptance temperature • Temperature stabilisation and dwell time for 1 hours • Verification of unit hot switching capability by switching off and on at maximum

operating acceptance temperature • \Unit cooling down in on mode to minimum operating acceptance temperature • Temperature stabilisation and dwell time for 1 hours • Unit heating up and cooling down to perform the required number of cycles • On last thermal vacuum cycle: unit performance test on maximum operating

acceptance temperature • on last thermal vacuum cycle: unit performance test on minimum operating

acceptance temperature • Unit heating up to ambient temperature in vacuum • Unit switch off at ambient temperature in vacuum • Ambient pressure recovery • Unit functional performance test at ambient (temperature and pressure) Furthermore, at all times where the unit is on, a selected set of parameters shall continuously be monitored (typically one measurement per minute).




T ambient

TNF max.

TFAmax.

TFAmin.

TNFmin.

TSUmin.

tS tS

tS tS tE

tE

P

Time

1

1

1

1

2

3

Unit on for corona test

Unit off

Unit on

2

tS

tS

4

Nomenclatur:

T ambient ambient temperature

TNFmax. maximum non-functioning temperature

TNFmin. miniumm non-functioning temperature

TFAmax. maximum acceptance temperature

TFAmin. minimum acceptance temperature

TSUmin. minimum start-up temperature

P pressure of TV chamber

tS dwell time = minimum stabilised temperature time = 1hour

tE minimum stabilised temperature time prior to start performance test:4 hours

1 performance test

2 unit on/off (verification of on/off switching capability at TSUmin.)

3 at all the times when unit is on, a selected set of major parameters shall continously

monitored (typically 1 point per minute)

4 for Corona test, some units are on during pressure drop

Figure 26 Acceptance Thermal Test Sequence

8.1.8 Proof Pressure Test The objective of the proof pressure test is: • to demonstrate that all pressurised tanks, sealed containers, lines, fittings, etc., can

withstand the maximum operating pressure including the qualification factors • to demonstrate that the leak rate is within the specified limits. The following requirements apply to items subjected to proof pressure qualification testing:

ENVREQ-829 : Test Set-up The test item shall consist of flight representative hardware and shall be in an operational mode which ensures that pressurisation/ vacuum of all compartments of the test item is representative of operation during flight.




The pressure shall be applied with an equivalent medium as specified to be used with the unit.

ENVREQ-830 : Test performance The following general test sequence shall be applied: • The pressure shall be applied in steps up to maximum operating pressure (limit

pressure) • The pressure shall be kept constant and the leakage shall be measured • The pressure shall be further increased in steps up to the proof pressure for flight unit

and up to the ultimate pressure for qualification test hardware where no rupture or collapse shall occur.

Since there are various standards applicable for designing and testing of pressurised vessels, the selected test method and the individual test pressure load values shall be defined in a test plan, which will be subject to space segment prime approval.

8.1.9 Leakage Test The purpose of the leakage test is to demonstrate the ability of a pressurised unit to meet the design leakage rate constraints specified in the unit specification.

ENVREQ-832 : The unit leak check shall be made prior to initiation of, and following the completion of unit qualification thermal and mechanical tests.

ENVREQ-833 : The leakage test shall be performed as follows: • The leakage tests shall be performed with unit pressurised at the maximum operating

pressure and then at the minimum operating pressure if the seals are dependent upon for proper sealing.

• The test duration shall be sufficient to detect any significant leakage.

8.1.10 Other Tests

8.1.10.1 Mechanism Tests

8.1.10.1.1 General This section describes the tests which shall be performed at mechanism level (and not at subassembly level).

ENVREQ-837 : The following rules apply: • Qualification of the mechanism will be obtained by the addition of a qualification test

sequence and a life test. Qualification shall be pronounced after the life test.




• Qualification cannot be obtained by proto-qualification (qualification shall be obtained by life test, and in consequence the qualification model cannot fly after the life test).

• When qualification is obtained, flight acceptance test sequence purpose is to verify workmanship and performances of the flight models.

All mechanism type shall be qualified on at least one unit. Qualification by similarity to a previous programme can be obtained only if all the following conditions are fulfilled: • Loads and worst case conditions are lower than previous mission on which the

mechanism has been qualified. • Mechanism is used with the same cycle (type, number, duration) • There is no modification on the hardware (and software if any). It is recommended to measure and control the friction torques at different stages in mechanism manufacturing. Verification of the performances of the mechanisms shall be possible on ground at unit level and assembly level. Verification of the operation shall be possible at system level. Use of special jigs or spacecraft attitude is acceptable.

8.1.10.1.2 Qualification tests

ENVREQ-839 : At least one mechanism assembly of each type shall be subjected to the following tests, with the assembly mounted in a representative manner, and the worst case loads applied and the worst case alignment.

i) The assembly shall be operated at ambient, and the performance measured (e.g. time to operate, resistance torque, minimum torque to operate, etc). The test results will be used to validate performances before and after environmental tests.

ii) In thermal vacuum conditions, the assembly shall be cooled to be at least 10 °C lower than the minimum expected temperature for flight operation, and the performance shall be measured.

iii) The assembly shall be heated in thermal vacuum to be at least 10 °C hotter than the maximum expected temperature for flight operation, and the performance measured.

iv) Worst case temperature gradient within the mechanism which could hamper the motion by increasing friction torques shall be represented at the maximum extent by the use of test heaters.

The torque margin shall be measured in ambient temperature (before and after EV tests), in hot and in cold environment for qualification. The mechanism drive unit (if the drive function is not part of the unit) used during the tests shall be electrically representative of the flight unit.

8.1.10.1.3 Acceptance tests The logic of tests to be performed at mechanisms level is identical to these defined for qualification tests. However the unit supplier can submit a proposal of re-adjusted test plan to be approved by the space segment prime.




ENVREQ-841 : During all tests the functional performances as well as the electrical characteristics shall be measured and compared to the qualification performances to verify that not significant changes have occurred. The torque margin shall be measured in ambient temperature (before and after EV tests), for acceptance.

8.1.10.1.4 Life tests

ENVREQ-843 : The following applicable tests shall be performed on any unit which experiences mechanical "wear-out" in operation. The orbit design life period shall be qualified in vacuum over a realistic temperature range and the specified number of cycles. The usual rule is to use the operating temperature range (for life test on thermal cycles) ; or the worst case temperature in the operating range (for mechanical life test) ; or a typical cycle in operating range representative of daily variation, all along the life test. The lifetime of the mechanisms shall be demonstrated / qualified by test in a configuration representative of the predicted worst case operational conditions with a minimum margin of 1.5 w.r.t. the mission utilization cycle. Before life test, the (QM or EQM) unit must have been submitted to qualification vibration levels. In any case, the life test conditions have to be approved by the space segment prime. The mechanism drive unit (if the drive function is not part of the unit) used during the tests shall be electrically representative of the flight unit. For deployment mechanisms, the subassembly which experiences the wear may be tested on its own, as a subassembly. The subassembly shall be loaded in a representative manner and operated the same number of times as designed. For rotating mechanisms, the subassembly which experiences the wear may be tested on its own, where this is possible (e.g. a single gyro from a unit containing other components). The unit or subassembly shall be loaded in a representative manner and operated under vacuum at temperature extremes, such that the total number of accumulated operations or cycles is as designed. For linear devices (e.g. electromagnetic valves, etc.), the subassembly which experiences wear may be tested on its own as a subassembly where this is possible, or as part of an assembly (e.g. thruster) or as a unit. The subassembly, unit or assembly shall be loaded in a representative manner and operated under vacuum at temperature extremes such that the total number of accumulated operations or cycles is as designed. Any unit or subassembly subject to gas or liquid throughput shall be tested under vacuum at temperature extremes such that the total test volume throughput is as designed. During such tests, the following conditions shall also be applied: a) The pressure (s) shall be generally representative of those expected in flight, but shall include a period equivalent to 10 % of volume throughput where the inlet pressure is higher than the maximum expected in flight, and a second period equivalent to 10 % of volume throughput, where the inlet pressure is lower than the minimum expected in flight. b) In the case of thrusters, each duty cycle expected during spacecraft life shall be tested such that the accumulated number for each duty cycle is as designed. Any element in a chain of actuation (motor, bearing, gear etc.) shall be compliant with the maximum number of cycles applicable to any of the remaining elements in the chain.




Lifetime of the critical mechanisms components is declared successful if the following is met: • No metal contact identified in the interface of solid lubricated surfaces. • Functional performances met with tolerances as specified for the entire life test • No rupture or loss of functionality of any part • Stiffness requirements met for the entire life test • Amount and size of wear product compliant with the contamination requirements

8.1.10.2 Solar Array Deployment Tests

ENVREQ-845 : The solar array qualification programme shall include a deployment test.

8.2 Satellite Level Tests Tests at satellite level are under the responsibility of the spacecraft contractor. It shall be proved that the overall spacecraft in its integrated condition is capable of withstanding the launcher flight environment based on a CLA performed by the launcher authority as well as the in-orbit conditions. This section specifies the test program which shall be conducted for the qualification and acceptance of the Galileo satellites.

ENVREQ-847 : Generally, the satellite shall be tested in a configuration as close as possible to the flight configuration as far as practicable. However, EQM units and/or QM units may be temporarily used on PFM/FM satellites as integration spares in case of unit malfunction. The provision of test connections on the satellite for performance measurements shall be such that, during all phases of assembly, integration and test of the satellite, the connection and disconnection of test equipment will not interfere with, or introduce a retrograde step in the assembly, integration and test, or invalidate the results of earlier performance measurements. For the Galileo IOV programme, the qualification approach is based on a STM (Structural Thermal Model) test programme followed by a PFM test programme.

8.2.1 Structural Thermal Model (STM) Tests

ENVREQ-850 : The pre- qualification of the satellite structure shall be performed with an STM which shall fulfil the following requirements: • The STM shall mechanically and thermally represent the spacecraft primary structure,

e.g. the STM structure is flight representative • The STM attachment mechanical interfaces shall be flight representative • The units shall be replaced by structural and thermal dummies which shall be defined

in the responsibility of the S/C contractor. The test programme for the qualification of the structure is defined in [RD 12].




8.2.1.1 Static Load Test The purpose of the static load test is to demonstrate by application of static or quasi-static loads that the spacecraft mechanical design is capable of sustaining the worst case launcher/dispenser induced static and dynamic accelerations without suffering permanent deformations or failure. The static load test is only performed on the STM and is not applicable to the integrated spacecraft.

ENVREQ-853 : The quasi- static acceleration levels to be applied shall be derived from the maximum quasi-static loads enveloping the results of the CLA results of the candidate launchers. If the mission profile indicates thermal stresses during the high acceleration periods, such stresses shall be determined analytically. A load equivalent to these thermal stresses shall be applied in addition to the above QSL. The STM test shall demonstrate the qualification factor specified in section 3.4.

8.2.1.2 STM Vibration Test

ENVREQ-855 : Modal survey test A modal survey shall be performed on the STM as defined in section 8.2.2.2. The following requirements apply to the satellite STM when subjected to vibration testing: • The STM/ SM shall be mounted to a rigid fixture via its standard flight interfaces • The vibration input shall be applied in each individual axis • The vibration levels (qualification levels with a sweep rate of 2 octave per minute) are

defined in section 4.1.2; they are derived from a launcher/dispenser/spacecraft CLA (including the respective qualification factors) and are submitted to the launcher authority approval.

• Where notching w.r.t. the specified levels is required, the proposed notching profile and the relevant justification analysis shall be submitted to the launcher authority for approval.

8.2.1.3 STM Shock Test The purpose of the STM shock test is to verify the satellite shock response when submitted to the launcher separation shock in order to confirm the validity of the shock environment specified to the units (cf. section 8.1.5.6 of the present document). This test is performed with the support of the launcher authority which provides the launcher interface and the flight representative separation system and which performs the separation release procedure.

ENVREQ-857 : Test mounting of the STM The STM shall be mounted on a flight representative launcher interface. It shall be hoisted and the separation system will be released. The shock generated at the satellite interface and within the satellite at the locations relevant to the satellite units shall be recorded.




8.2.1.4 STM Acoustic Test The purpose of the STM acoustic test is to verify the capability of the structure to be submitted to the qualification acoustic level as well as to perform an early verification of the satellite acoustic response in order to confirm the validity of the random environment specified to the units (cf. section 8.1.5.4 of the present document).

ENVREQ-859 : Acoustic Qualification The STM shall be submitted to the qualification acoustic level. The response of the STM shall be recorded. The STM shall be submitted to inspection after test.

8.2.1.5 STM Thermal Balance Test The purpose of the STM Thermal Balance Test is to verify early in the satellite development programme the satellite thermal behaviour and the ability of the thermal control design to meet the required performances.

ENVREQ-861 : Equipment wit dummies The STM shall be equipped with thermal dummy units, which adequately represent the thermal behaviour of the flight units (in terms of conductive and radiative heat transfer) and adequate heaters to be representative of the thermal dissipation of the units. In addition the thermal control subsystem hardware shall be flight representative. In general, the purpose of the thermal balance testing is to demonstrate the ability of the thermal control subsystem to maintain temperatures inside the specified design temperatures and to verify that thermal control performs correctly under vacuum and thermal conditions expected to be encountered during the mission. Its purpose is also to verify the Thermal Mathematical Model of the spacecraft. The following requirements apply:

ENVREQ-864 : General requirements The definition of the test cases shall take into account the flight conditions as far as practical.

ENVREQ-865 : Test set-up The thermal state of the spacecraft should be obtained by: • simulating the incident radiation (using radiation sources simulating the solar and

albedo radiation intensities), • simulating the heat absorbed by the spacecraft with high accuracy (with help of

heater blankets and infrared lamps) or combining these two methods. If necessary internal heaters shall be defined and implemented by the satellite test contractor.

The location of the necessary thermo-couples will be defined by the thermal control subsystem and the space segment prime. The spacecraft shall be operated under vacuum conditions in thermal environments similar to those expected during orbital life. The TB test thermal vacuum environment shall have a pressure of 1E-5 hPa or less.




The baseline orientation of the spacecraft in the vacuum test chamber is such that the Y panels are horizontal (for correct operation of the 2D heat pipe network located on the Y panels), with +Y or -Y oriented upwards (TBD in the TV test specification). On all external surfaces flight representative thermal hardware shall be applied (MLI, OSR etc.). A test model of the SA Yoke shall be defined by the satellite test contractor depending on the objective of the test.

ENVREQ-866 : Test conductance The temperature limits of the TB test will be derived from the results of the overall S/C thermal analysis and will be detailed in the TB test S/C Thermal Test Requirements. The TB test shall include thermal balance test phases as well as test phases where the performance of the thermal regulation loops are checked (especially those controlling the clocks). Sufficient temperature sensors shall be installed on and in the spacecraft to provide all necessary information to verify the thermal analysis. Success criteria for the TB test shall be: • a demonstration of satisfactory operation of the S/C within the specified temperature

values, • a satisfactory degree of accuracy of test parameter accuracy (power dissipation,

temperatures, absorbed fluxes) to allow accurate TMM correlation.

8.2.1.6 Micro-Vibration Test The purpose of the micro-vibration test is to characterise the micro-vibration environment which applies on the satellite in order to verify the micro-vibration specification applying on the units. This test will be performed by operating a reaction wheel and by measuring the generated micro-vibration levels at the locations where micro-vibration sensitive units are located.

ENVREQ-868 : Conductance of the test The STM shall be submitted to a micro-vibration test. A reaction wheel model shall be implemented on the STM (instead of the reaction wheel dummy unit). The STM shall be instrumented at the locations where sensitive units are located. The reaction wheel shall be operated (spin-up, steady-state and spin-down) and the micro-vibration environment shall be measured and recorded.

8.2.1.7 Physical Properties The purpose of the physical properties measurement performed on the STM is to determine the STM mass, centre of gravity location, moments of inertia around its three co-ordinate axes.

ENVREQ-870 : Test configuration This test shall be performed with the STM in a configuration as close as possible to the flight models configuration. Any deviation w.r.t. the flight models configuration shall be carefully recorded in order to allow the flight model configuration physical properties to be predicted based on the obtained measurements. For that purpose, all items shall be weighted with the required accuracy and their locations shall be clearly and precisely defined.




8.2.2 Structural Model (SM) Tests

ENVREQ-872 : The following tests shall be performed on the SM: • Physical properties • Sine vibration test

ENVREQ-873 : The physical properties test to be performed on the SM shall be as specified in section 8.2.1.7.

ENVREQ-874 : The sine vibration test to be performed on the SM shall be as specified in section 8.2.1.2 except for the levels and sweep rate which shall be the acceptance ones.

8.2.3 Proto-flight Model Satellite Test

8.2.3.1 General

ENVREQ-877 : The following requirements shall apply: For the first built flight model of the Galileo spacecraft the Proto-flight qualification approach shall be applied for satellite design qualification. The PFM satellite shall be a standard flight model. It shall be equipped with qualified proto-flight and flight units and it shall be capable of being launched and of meeting its design requirements for the operational life time. The applied test program is considered to be qualification and acceptance at the same time.

8.2.3.2 Proto-flight Test Program

8.2.3.2.1 Mechanical Tests

8.2.3.2.1.1 Modal Survey The purpose of the modal survey is to determine experimentally the natural frequencies and damping factors of the S/C throughout the dynamically relevant range.

ENVREQ-882 : Experimental determination of the natural frequencies, mode shapes, and damping factors of the S/C shall be accomplished by exciting the natural modes either by single-point or multi-point excitation with transient, random or sinusoidal characteristics and with accelerometers strategically distributed throughout the test article to monitor the response. The results of the modal survey shall be used: • to check whether any natural frequency falls into an undesirable range,




• to check or adjust the natural frequencies and modes determined analytically with the help of a finite element mathematical model of the structural system,

• to have a reference to be compared to after application of the qualification levels Once the mathematical model is in agreement with the physical system, it can be used to predict structural responses such as displacements, accelerations, and stresses for any force input for the structural system alone, or, if the adjacent element is compatibly represented, for combined structural system. Superimposing of thermal stresses can also be performed.

8.2.3.2.1.2 Sine Vibration Test The purpose of the sine vibration test is to demonstrate that the satellite withstands the vibration environment encountered during launch or other high vibration exposures and defined in section 4.1.1.2.2.

ENVREQ-885 : The sine vibration levels shall be the qualification levels with a sweep rate of 4 octaves per minute.

ENVREQ-886 : The satellite shall be placed in the same functional configuration as intended for launch. All units which shall be on during launch shall be switched on, placed in the relevant functional operating mode and shall be monitored during the test.

8.2.3.2.1.3 Acoustic Test The purpose of the acoustic test is to demonstrate the ability of the spacecraft and the unit to withstand the qualification level acoustic environment of the launch.

ENVREQ-888 : The satellite shall be placed in the same functional configuration as intended for launch. All units which shall be on during launch shall be switched on, placed in the relevant functional operating mode and shall be monitored during the test.

ENVREQ-889 : The qualification level sound pressure levels are presented in section 4.1.2. They correspond to an envelope of the acoustic spectra of the candidate launchers and allow therefore proto-qualification w.r.t. the acoustic environment of all applicable launchers to be obtained. This corresponds to the baseline approach. However it must be noted that this leads to apply on the PFM satellite acoustic levels which are more severe than the qualification levels strictly required by all individual launchers (since acoustic spectrum of some launchers are overlapping). Therefore, if some issues are detected during the STM acoustic test (e.g. unexpected high response in given frequency bands), a less conservative approach could be proposed and discussed with the launcher authorities instead such as:




• Perform on the IOV PFM satellite the proto-qualification acoustic spectrum covering Soyuz only and using a PFM approach for the first FOC satellite using a different launcher

• Submit to the IOV PFM satellite for a test duration of 30 s the proto-qualification acoustic spectrum covering Soyuz and then to the proto-qualification acoustic spectrum of Ariane 5 for another test duration of 30 s.

• Submit on one IOV satellite the proto-qualification acoustic spectrum covering Soyuz and on another IOV satellite the proto-qualification acoustic spectrum covering Ariane 5 for instance

ENVREQ-895 : Applicable test tolerances The table below presents the applicable test tolerances (taking into account standard test equipment inaccuracy):

Octave band Frequency (Hz)

Test Tolerance dB

25-40 -2, +4 63-2000 -1,+3

Table 113 - Applicable test tolerances

8.2.3.2.1.4 PFM Shock Test The purpose of the shock test performed on the PFM is to demonstrate that the spacecraft in the applicable configuration corresponding to the relevant shock event, withstands the shock levels and frequency spectra as predicted for flight without any operational anomalies. The spacecraft is submitted to the two following main shock events: • The launcher separation shock • The solar array pyro release shock

ENVREQ-899 : Launcher separation shock test The PFM shock test shall be performed in a similar manner as for the STM shock test. The satellite shall be instrumented with accelerometers at the satellite to launcher interface and at the locations relevant for the shock environment. However, in this case, the satellite shall be placed in the same functional configuration as intended for launch. All units which shall be on during launch shall be switched on, placed in the relevant functional operating mode and shall be monitored during the test. After application of the shock (the satellite shock response being recorded), it shall be checked that no modification of the satellite functional configuration has occurred and that no anomaly has been detected.

ENVREQ-900 : Solar array pyro release The solar array pyro release and deployment test shall be performed at satellite level after the satellite mechanical tests. The satellite shall be instrumented with accelerometers at the solar array interface and at the locations relevant for the shock environment. The satellite shall be placed in the same functional configuration as intended for the solar array




deployment performed in flight. All units which shall be on in this phase shall be switched on, placed in the relevant functional operating mode and shall be monitored during the test. After application of the shock (the satellite shock response being recorded), it shall be checked that no modification of the satellite functional configuration has occurred and that no anomaly has been detected.

8.2.3.2.1.5 Pressure Test The purpose of pressure tests on the spacecraft is to demonstrate that the integrated pressurised subsystems achieve the specified requirements for pressure integrity and leakage rate.

ENVREQ-902 : Pressure Test For that purpose, the pressurised elements of the satellite shall be placed at the MEOP.

8.2.3.2.1.6 Leakage Test The purpose of the satellite leakage test is to demonstrate that all pressurised units when pressurised as defined in the previous section are still leak tight after having been submitted to the satellite environmental test programme.

ENVREQ-904 : Leakage Test After having pressurised at their MEOP all satellite pressurised units, the satellite overall leakage rate shall be measured and shall be controlled w.r.t. the overall leakage requirement.

8.2.3.2.1.7 Physical Properties The purpose of the physical properties measurement is to determine the spacecraft mass, centre of gravity location, moments of inertia around its three co-ordinate axes.

ENVREQ-906 : Physical properties measurement This test shall be performed with the satellite in a configuration as close as possible to the flight configuration. Any deviation w.r.t. the flight configuration (presence of non-flight items or missing flight items) shall be carefully recorded in order to allow the flight configuration physical properties to be predicted based on the obtained measurements. For that purpose, all non-flight items shall be weighted with the required accuracy and their locations shall be clearly and precisely defined.

8.2.3.2.2 Thermal Vacuum (TV) and Thermal Balance (TB) Tests The following requirements apply to TV/ TB testing:




ENVREQ-908 : General requirements TV and TB test shall be combined if possible. The definition of test cases shall take into account all possible spacecraft internal redundancy configurations.

ENVREQ-909 : Test set-up The baseline orientation of the spacecraft in the vacuum test chamber is identical to the one used for the STM TB test. On all external surfaces flight representative thermal hardware shall be applied (MLI, OSR etc.).

8.2.3.2.2.1 Thermal Balance Test (TB Test) The purpose of the PFM thermal balance test is to validate the satellite thermal behaviour (correlation of the satellite Thermal Mathematical Model), to demonstrate the ability of the thermal control subsystem to maintain temperatures inside the specified design temperatures and to verify that thermal control performs correctly under vacuum and thermal conditions expected to be encountered during the mission. This test allows the Thermal Control Subsystem to be formally qualified.

8.2.3.2.2.2 PFM Satellite Thermal Vacuum Test (TV) The purpose of the PFM satellite TV test is to demonstrate the ability of the satellite to operate properly under the thermal vacuum environment that simulates the worst in-orbit conditions (including the qualification margin). The following requirements shall apply for the satellite qualification model level test for the PFM TV test:

ENVREQ-912 : The maximum temperatures shall be selected such that the spacecraft units are at their maximum temperature flight predicted limits increased by the qualification temperature margin. The minimum temperature shall be selected such that the spacecraft units are at the minimum flight predicted temperature limit decreased by the qualification margin. For proto-qualification testing, at least three (3) thermal cycles in vacuum shall be completed. The configuration of the spacecraft shall be defined in the related Satellite TV -Test Procedure. Functional tests shall be performed at each of the extreme temperature plateaus with a continuous monitoring of all necessary outputs. The duration of a plateau should be not less than two (2) hours after unit temperature stabilisation. During transitions between plateaus, adequate monitoring of the satellite units shall be performed. Initial ambient and final ambient performance tests shall be also carried out in order to check that no degradation of the performances can be detected as a result of the thermal vacuum cycling test.




The thermal vacuum test shall be preceded by the depressurisation test which objective is to demonstrate that the satellite can be submitted to the depressurisation without any degradation nor anomalies (e.g. Corona). The satellite is placed in the same functional configuration as intended for launch. All units which shall be on during launch shall be switched on, placed in the relevant functional operating mode and shall be monitored during the test. At the end of the PFM satellite PFM test, the satellite shall be inspected and it shall be verified for instance that no damage of the satellite MLI can be observed as a result of the depressurisation phase. Furthermore, contamination samples shall be placed in the vacuum chamber and these samples shall be analysed.

8.2.3.2.3 Solar Array Deployment Test

ENVREQ-914 : A solar array deployment test shall be performed at PFM level after the satellite mechanical tests. It shall be performed using flight standard actuators, e.g. pyro-actuators if applicable. Furthermore, the same release sequence as intended for the flight shall be used for the test. Refer also to section 8.2.2.2.1.4.

8.2.4 Spacecraft Acceptance Test Program

ENVREQ-916 : Spacecraft Acceptance Test Program The following test program shall be submitted to each recurrent spacecraft: • Sinusoidal Vibration Test • Acoustic Noise Test (or alternative TBD) • Thermal Vacuum Test • Pressure and Leak Test • Solar Array Deployment Test • Physical Properties

8.2.4.1 Sinusoidal Vibration Test

ENVREQ-918 : Sinusoidal Vibration Test A sine vibration tests shall be performed at acceptance level in order to detect latent materials or detect workmanship defects, and to compare the resonance frequency distribution with that of the mathematical model or modal survey and with the previous satellite flight models. The test is identical to the PFM test except for the applied levels which shall be acceptance levels. Furthermore since the IOV satellites are only intended for launch on Soyuz, the acceptance sine levels applicable to Soyuz may be used in order to minimize the stress of the flight hardware.

8.2.4.2 Acoustic Test The purpose of the acoustic test is to detect material and workmanship defects that are not detected in a static condition or by the sine test.




ENVREQ-920 : Acoustic Test The acceptance level sound pressure levels and test duration are presented in section AD. The test tolerances are identical to the PFM acoustic test. The test is identical to the PFM test except for the applied levels which shall be acceptance levels. Furthermore since the IOV satellites are only intended for launch on Soyuz, the acceptance sine acoustic levels applicable to Soyuz may be used in order to minimize the stress of the flight hardware.

8.2.4.3 Pressure and leakage

ENVREQ-922 : Pressure and leakage tests For acceptance pressure tests, the subsystems of the spacecraft shall be pressurised to the maximum expected operating pressures and kept at this pressure for a sufficient duration to establish that the leak rates are within the specified limits.

8.2.4.4 Acceptance Thermal Vacuum Tests

ENVREQ-924 : Acceptance Thermal Vacuum Tests The acceptance thermal vacuum test performed on the FM is identical to the PFM thermal vacuum test except on the following aspects: • no thermal balance test is required • the acceptance temperature extremes (maximum flight predicted temperatures

increased by the acceptance margin and minimum flight predicted temperatures decreased by the acceptance margins) shall be applied on the satellite units

8.2.4.5 Solar Array Deployment Test

ENVREQ-926 : Solar Array Deployment Test This test is identical to the test performed on the satellite PFM except that no shock instrumentation is needed.

8.2.4.6 Physical Properties

ENVREQ-928 : Physical Properties tests A physical properties test shall be performed on each recurrent spacecraft. This test shall be performed as specified in section 8.2.3.2.1.7.