Auxiliary Passengers using Arianespace Systems

User’s Manual Issue 1 – Revision 0 June 2017

Issued and approved by Arianespace

Roland Lagier

Chief Technical Officer

Auxiliary Passengers User’s Manual Issue 1

Preface

The present Auxiliary Passengers User’s Manual provides an overview of the Arianespace launch service using the systems for auxiliary payload available on Soyuz and Vega launchers, which together with Ariane 5 constitute the European space transportation union operated by Arianespace at the Guiana Space Centre. This document contains the essential data which are necessary: ∑ To assess compatibility of micro and mini spacecraft mission with available systems, ∑ To initiate the preparation of all technical and operational documentation related to a launch of spacecraft as auxiliary passenger on a Soyuz or Vega mission. Further information regarding Soyuz and Vega launch systems can be found respectively in the Soyuz and Vega User’s Manual. Inquiries concerning clarification or interpretation of this manual should be directed to the addresses listed below. Comments and suggestions on all aspects of this manual are encouraged and appreciated.

France Headquarters USA - U.S. Subsidiary Arianespace Arianespace Inc. Boulevard de l'Europe 601 13th Street N.W. Suite 710 N. BP 177 91006 Evry-Courcouronnes Cedex France Washington, DC 20005, USA Tel: +(33) 1 60 87 60 00 Tel: +(1) 202 628-3936 Fax: +(33) 1 60 87 64 59 Fax: +(1) 202 628-3949

Singapore - Asean Subsidiary Japan - Tokyo Office Arianespace Singapore PTE LTD Arianespace # 18-09A Shenton House Kasumigaseki Building, 31Fl. 3 Shenton Way 3-2-5 Kasumigaseki Chiyoda-ku Singapore 068805 Tokyo 100-6031 Japan Fax: +(65) 62 23 42 68 Fax: +(81) 3 3592 2768

Website French Guiana - Launch Facilities Arianespace BP 809 www.arianespace.com 97388 Kourou Cedex French Guiana Fax: +(33) 5 94 33 62 66

An updated version of the present User’s Manual will be subsequently released introducing additional small satellites launch opportunities with VEGA-C and ARIANE A6 launch systems.

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Configuration Control Sheet

Issue / Prepared Approved Date Change description Rev. by by Draft May, 2008 First issue Issue 1 June, 2017 Signed issue C. DUPUIS J. THIERY Revision 0

IV Arianespace ©, June 2017 Auxiliary Passengers User’s Manual Issue 1 Table of contents Preface Configuration control sheet Table of contents Acronyms, abbreviations and definitions

CHAPTER 1. INTRODUCTION 1.1. PURPOSE OF THE USER’S MANUAL FOR AUXILIARY PASSENGERS USING ARIANESPACE SYSTEMS

1.2. AUXILIARY PASSENGERS DEFINITION 1.2.1. General 1.2.2. Auxiliary Passengers classification

1.3. PROCURING AUXILIARY PASSENGER LAUNCH SERVICE

1.4. FLIGHT OPPORTUNITIES FOR AUXILIARY PASSENGERS

1.5. SPECIFIC RULES ASSOCIATED TO AUXILIARY PASSENGERS 1.5.1. General considerations 1.5.2. Rules applicable to auxiliary passengers

1.6. ARIANESPACE SYSTEMS FOR AUXILIARY PASSENGERS 1.6.1. Introduction 1.6.2. ASAP-S 1.6.3. VESPA

CHAPTER 2. INTERFACES FOR AUXILIARY PASSENGERS 2.1. GENERAL

2.2. REFERENCE AXES

2.3. ENCAPSULATED SPACECRAFT INTERFACES 2.3.1. General 2.3.2. Payload usable volume definition 2.3.3. Special on-fairing insignia

2.4. MECHANICAL INTERFACE

2.5. ELECTRICAL INTERFACES 2.5.1. Lines definition 2.5.2. Spacecraft to EGSE umbilical lines 2.5.3. L/V to spacecraft electrical functions 2.5.4. Electrical continuity interface

2.6. INTERFACES VERIFICATIONS 2.6.1. Prior to the launch campaign 2.6.2. Pre-launch validation of the electrical I/F

CHAPTER 3. AUXILIARY PASSENGERS LAUNCH MISSION

3.1. INTRODUCTION 3.2. TYPICAL MISSION PROFILE 3.3. MISSION DURATION

Arianespace ©, June 2017 V Auxiliary Passengers User’s Manual Issue 1 3.4. SPACECRAFT ORIENTATION DURING THE ASCENT PHASE 3.5. SEPARATION CONDITIONS 3.5.1. Orientation performance 3.5.2. Separation mode and pointing accuracy

CHAPTER 4. MICRO AUXILIARY PASSENGER DESIGN AND VERIFICATION REQUIREMENTS 4.1. INTRODUCTION

4.2. DESIGN REQUIREMENTS FOR MICRO AUXILIARY PASSENGER 4.2.1. Safety Requirements 4.2.2. Selection of Spacecraft Materials 4.2.3. Micro S/C Mass Properties 4.2.4. Frequency Requirements for Micro Auxiliary Passenger 4.2.5. Design Loads 4.2.6. Line loads peaking on the spacecraft 4.2.7. Line loads peaking induced by the Micro Auxiliary Passenger 4.2.8. Handling Loads during ground operations 4.2.9. Local Loads 4.2.10. Dynamic Loads

4.3. MICRO SATELLITES COMPATIBILITY VERIFICATION REQUIREMENTS 4.3.1. Verification logic 4.3.2. Safety factors 4.3.3. Spacecraft compatibility tests

4.4. THERMAL LOADS

4.5. RF ENVIRONMENT

CHAPTER 5. MINI AUXILIARY PASSENGER DESIGN AND VERIFICATION REQUIREMENTS 5.1. INTRODUCTION

5.2. DESIGN REQUIREMENTS FOR MINI AUXILIARY PASSENGER 5.2.1. Safety Requirements 5.2.2. Selection of Spacecraft Materials 5.2.3. Mini S/C Mass Properties 5.2.4. Frequency Requirements for Mini Auxiliary Passenger 5.2.5. Design Loads 5.2.6. Line loads peaking on the spacecraft 5.2.7. Line loads peaking induced by the Mini Auxiliary Passenger 5.2.8. Handling Loads during ground operations 5.2.9. Local Loads 5.2.10. Dynamic Loads

5.3. MINI SATELLITES COMPATIBILITY VERIFICATION REQUIREMENTS 5.3.1. Verification logic 5.3.2. Safety factors 5.3.3. Spacecraft compatibility tests

5.4. THERMAL LOADS

5.5. RF ENVIRONMENT

§1 - p6 Arianespace ©, June 2017 Auxiliary Passengers User’s Manual Issue 1 CHAPTER 6. MISSION MANAGEMENT & LAUNCH CAMPAIGN ORGANISATION FOR AUXILIARY PASSENGERS 6.1. INTRODUCTION

6.2. MISSION MANAGEMENT 6.2.1. Contract organization 6.2.2. Schedule 6.2.3. Meetings and Reviews

6.3. SYSTEMS ENGINEERING SUPPORT 6.3.1. Interface Management 6.3.2. Mission Analysis 6.3.3. Auxiliary Passenger Compatibility Verification 6.3.4. Post-launch Analysis

6.4. LAUNCH VEHICLE ADAPTATION

6.5. LAUNCH CAMPAIGN 6.5.1. Typical Auxiliary Passenger launch campaign 6.5.2. Summary of launch campaign meetings and reviews 6.5.3. Range Support

6.6. SAFETY ASSURANCE

6.7. QUALITY ASSURANCE

6.8. OPTIONAL SERVICES

Annex 1 – APPLICATION TO USE ARIANESPACE’S LAUNCH VEHICLE (DUA) TEMPLATE

Annex 2 – STANDARD AUXILIARY PASSENGER ADAPTERS 2.1 SSASAP5 ring 2.2 PAS 381 S adapter 2.3 PAS 432 S adapter 2.4 PAS 610 S adapter 2.5 AR 937 adapter Annex 3 – ARIANESPACE AUXILIARY PASSENGERS LAUNCH RECORD

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Acronyms, abbreviations and definitions

A ACS Attitude Control System AE Arianespace ASAP Arianespace System for Auxiliary Payloads B BT POC Combined operations readiness Bilan Technique Plan d’ Opérations review Combinées C CAD Computer Aided Design CBOD-LT Low Tension Clamp Band Opening Device CDR Critical Design Review CFRP Carbon Fiber Reinforced Plastic CLA Coupled Loads Analysis CM Mission Director Chef de Mission CNES French National Space Agency Centre National d’ Etudes Spatiales CoG Center of Gravity COTE Check-Out Terminal Equipment CP Program director Chef de Programme CRAL Post Flight Debriefing Compte-Rendu Après Lancement CSG Guiana Space Centre Centre Spatial Guyanais CVCM Collected Volatile Condensable Material CVI Real time flight evaluation Contrôle Visuel Immédiat D DAMF Final mission analysis document Document d' Analyse de Mission Fi nale DAMP Preliminary mission analysis Document d' Analyse de Mission document Préliminaire DCI Interface control document Document de Contrôle d’ Interface DUA Application to use Arianespace Demande d' Utilisation Arianespace launch vehicles E EGSE Electrical Ground Support Equipment EIRP Equivalent Isotropic Radiated Power ELV European Launch Vehicle S.p.A. EMC Electro magnetic Compatibility EPCU Payload preparation complex Ensemble de Préparation des Charges Utiles ESA European Space Agency F FAR Flight Acceptance Review FEM Finite Element Model FM Flight Model FQR Final Qualification Review FSA Russian Federal Space Agency

§1 - p8 Arianespace ©, June 2017 Auxiliary Passengers User’s Manual Issue 1 G GRS General Range Support GSE Ground Support Equipment H HPF Hazardous Processing Facility I Isp Specific impulse ITAR International Traffic in Arms Regulations K KRU Kourou L LAM Measuring instrument laboratory La boratoire Mesures LBC Check out equipment room Laboratoire Banc de Contrôle LEO Low Earth Orbit LEOP Launch and Early Orbit Phase LPSS Launcher Payload Separation System LSA Launch Service Agreement LTAN Local Time of Ascending Node LTDN Local Time of Descending Node LV Launch Vehicle LW Launch Window M MCI Mass, balances and inertias Masse, Centre de gravité, Inerties MGSE Mechanical Ground Support Equipment MLI Multi Layer Insulation MMH Mono methyl Hydrazine MUS Soyuz at CSG User's Manual Manuel Utilisateur Soyuz du CS G MUV Vega User's Manual Manuel Utilisateur Vega N N/A Not Applicable O OASPL Overall Acoustic Sound Pressure Level OCOE Overall Check Out Equipment P PAF Payload Attachment Fitting PAS Payload Adapter System PDR Preliminary Design Review PFM Proto-Flight Model POC Combined operations plan Plan d’ Opérations Combinées POI Interleaved Spacecraft Operations Plan d’ Opérations Imbriquées Plan POS Spacecraft operations plan Plan d’Opérations Satellite PPF Payload Preparation Facility ppm parts per million PSD Power Spectral Density Q QA Quality Assurance QR Qualification Review QSL Quasi-Static Load

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R RAAN Right Ascension of the Ascending Node RAL Launch readiness review Revue d’ Aptitude au Lancement RAMF Final mission analysis review Revue d' Analyse de Mission Finale RAMP Preliminary mission analysis review Revue d' Analyse de Mission Préliminaire RAV Launch vehicle flight readiness Revue d’ Aptitude au Vol du lanceur review RF Radio Frequency RMS Root Mean Square rpm revolutions per minute RPS Spacecraft preparation manager Responsable Préparation Satellite RS Safety manager Responsable Sauvegarde RSG Ground safety officer Responsable Sauvegarde Sol RSV Flight safety officer Responsable Sauvegarde Vol S S/C Space craft SLV Vega Launch Site Site de Lancement Vega SOW Statement Of Work SRS Shock Response Spectrum SSO Sun Synchronous Orbit STM Structural Test Model T TBC To Be Confirmed TBD To Be Defined TC Tele command TM Tele metry U UC Upper Composite UCIF Upper Composite Integration Facility UT Universal Time V VESPA VE ga Secondary Payload Adapter

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INTRODUCTION Chapter 1

1.1 Purpose of the Auxiliary Payload User’s Manual

Arianespace has been launching Auxiliary Passengers since the early days of Ariane in 1980. With the introduction of the Ariane Structure for Auxiliary Payloads on the Ariane 4 and Ariane 5 family of vehicles, Arianespace initiated a standardized approach. This allowed many teams from the worldwide educational, scientific, military and industrial communities to gain easy and cost effective access to space for their small projects. Thanks to the Vega and Soyuz launch systems now fully operational at the Guiana Space Centre, Arianespace continues to propose ride share opportunities on larger passenger mission. Several flights of Soyuz (VS02, VS14) and Vega (VV02, VV07), have demonstrated the small sat capabilities using dedicated carrying structures for auxiliary passenger (respectively the Arianespace Structure for Auxiliary Payloads for Soyuz - ASAP-S and the VEga Secondary Payload Adapter - VESPA). This makes the life always easier to the companies, agencies, universities or other entities that want to take advantage of affordable Auxiliary Passenger system for access to space.

The present User’s Manual for Auxiliary Passengers is intended to provide basic information for Auxiliary Passengers on the Arianespace’s launch service solutions using the Soyuz or Vega launch systems operated from the Guiana Space Centre. Furthermore, it defines the rights and obligations that apply to Auxiliary passengers. The content encompasses: • the definition of an Auxiliary Passenger; • the ASAP-S and VESPA description; • the description of the interfaces between small Spacecraft and Launch Vehicle (LV); • the requirements for small Spacecraft design and verification; • the mission integration and management for Auxiliary Passenger; • the processing and ground operations performed at the launch site for Auxiliary Passenger.

An updated version of the present User’s Manual will be subsequently released introducing additional small satellites launch opportunities with VEGA-C and ARIANE A6 launch systems. Together with the Soyuz from CSG User’s Manual, the Vega User’s Manual, the Payload Preparation Complex Manual (EPCU User’s Manual) and the CSG Safety Regulations, it gives readers the information to assess the compatibility with the proposed standardized configurations. Dedicated configurations can be implemented as well on a case by case basis. For more detailed information, the reader is encouraged to contact Arianespace.

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1.2 Auxiliary passenger definition 1.2.1. General An Auxiliary Passenger Customer takes advantage of extra performance and volume available on a Main Passenger mission either on Soyuz or Vega. The main characteristics of the mission are defined by the Main Passenger (or co-passengers), including launch period and launch date, launch time (or window), targeted orbit, etc ….

1.2.2. Auxiliary passenger classification The proposed configuration and the associated Launch Service depend on the mass and volume of the Auxiliary Passenger.

Arianespace has established an Auxiliary Passenger classification which corresponds to the available positions on the existing multiple launch systems (ASAP-S and VESPA, refer to chapter 1.6 for a detailed description).

Main passenger

Typically Main Main 2500 kg passenger passenger

Typically Typically 800 kg 800 kg

Mini Mini Micro Micro Micro Micro Micro Micro

The following small sat classification applies in the present User’s Manual: • Mini : small satellite with a mass between 200 and 400 Kg, compatible in volume with the central ASAP-S position or the internal VESPA position. • Micro : small satellite with a mass between 50 and 200 Kg, compatible in volume with the external ASAP-S position or half of the internal VESPA position.

The available accommodations are summarized in the table below:

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Micro S/C Mini S/C 50 to 200 Kg 200 to 400 Kg SOYUZ / ASAP-S External positions 4 SOYUZ / ASAP-S Central position 1, 2 or 3 1 VEGA / VESPA Internal position 1 or 2 1

Table 1.2.1: Available options for small sat

For smaller spacecraft with a mass below 50 kg and for canisterized Cubesat (either 1U, 3U or 6U), please contact Arianespace.

1.3 Procuring Auxiliary Passenger Launch Services

A template of the Application to Use Arianespace’s Launch Vehicle for Auxiliary Passengers is available in Annex 1.

It is used by Arianespace to identify the launch opportunities and provide preliminary mission and pricing information.

The Launch Services Agreement (LSA) can be signed when the launch opportunity(ies) is(are) identified, or at an earlier stage, with the advantage to get priority on the opportunities to come.

The mission integration process for Auxiliary Passenger is fully described in chapter 6 of the present manual.

1.4 Launch opportunities for Auxiliary Passengers

The launch of an auxiliary passenger is linked to the existence of a launch opportunity i.e. a mission for a main passenger where some extra volume and performance capabilities are available.

Arianespace is committed to identify and maintain such launch opportunities for auxiliary payload. The corresponding missions are mainly for Sun synchronous Orbit (SSO) and Low Earth Orbit (LEO), however some opportunities also arise on Geostationary Transfer Orbit (GTO) or other orbits.

Arianespace maintains a list of the missions which are suitable for auxiliary passengers on Soyuz and Vega launchers with the corresponding launch period.

1.5 Specific rules associated to Auxiliary Passengers 1.5.1 General considerations The Auxiliary Passenger is part of a launch dedicated to a main passenger therefore it has to adapt to the main passenger schedule and mission requirements.

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1.5.2 Rules applicable to auxiliary passengers

1.5.2.1 Main passenger approval

The Launch of an auxiliary passenger is subject to the approval of the main passenger. The auxiliary passenger shall be able to answer to any inquiry aiming at demonstrating innocuousness for main passenger.

1.5.2.2 Schedule

The Launch of an auxiliary passenger is totally subordinated in terms of schedule to the launch schedule of the main passenger. It implies that an auxiliary passenger shall in no case be entitled to affect the launch schedule. The consequence is that, should an Auxiliary Payload not be ready for the Launch, it will fly as it is if already mounted on the upper composite or, it will be replaced by a dummy payload.

1.5.2.3 Dummy payload

For all auxiliary passenger, a dummy payload has to be made available in case the actual spacecraft is not ready for the launch.

The dummy payload provided by the Customer must be representative of its satellite in terms of mass, center of gravity and mechanical interface. The volume of a dummy payload must be smaller (or identical) to the actual satellite volume. No electrical interface is required.

The dummy payload must be compatible of the flight environment and mission.

Proof of availability of the dummy will be made by Auxiliary Passenger Customer for or prior to the Flight Readiness Review (RAV) and the dummy shall be provided at the beginning of the launch campaign.

1.5.2.4 Targeted orbit The targeted orbit is defined following the main passenger technical requirements. In particular, for Sun synchronous Orbit (SSO) mission, the local time of ascending node (LTAN) is defined by the main passenger. For Mini Auxiliary Passenger in central ASAP-S position or internal VESPA position, the re- ignition capabilities of the Vega and Soyuz upper stages allow, at a certain extent depending on available performance margins, some modification of the orbit altitude, eccentricity and inclination to attain the specific needs of the Mini Auxiliary Passenger. For Micro Auxiliary Passenger and Cubesats, the obtained orbit will be defined by Arianespace.

1.5.2.5 Flight sequence and separation

The mission profile, separation conditions and pointing and separation time are defined by Arianespace. The auxiliary passenger shall be compatible with the sizing scenario in terms of flight duration, angular velocity at separation, etc …

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For Mini Auxiliary Passenger in ASAP-S central position or VESPA internal position, and depending on overall mission, the customer preferred separation conditions can be considered.

1.5.2.6 Auxiliary Passenger status during final countdown and flight

The Auxiliary Passenger shall be inert (S/C OFF, no RF emission, no status changes) during the final countdown and ascent phase up to TBD minutes after the auxiliary passenger separation. The actual delay after separation will be determined by Arianespace depending on the type of orbit, mission timeline and main passenger requirements.

1.5.2.7 Auxiliary Passenger preparation at the launch site

The Auxiliary Passenger shall be ready the day before the start of the Combined Operations (POC) for integration on its adapter, and then on the carrying structure. No access to the satellite is authorized after mating on the carrying structure.

The launch campaign for auxiliary passenger is fully described in chapter 6.

1.6 The Arianespace carrying systems for auxiliary passengers

1.6.1. Introduction In order to provide launch opportunities to small satellites, Arianespace has developed dedicated carrying systems to carry and deploy small satellites. • On Soyuz, the so-called Arianespace Structure for Auxiliary Payloads for Soyuz (ASAP-S) allows to embark up to four (4) micro satellites on the four external positions together with one (1) mini satellite in central position or up to three (3) micro satellites in central position. • On Vega, the so-called VEga Secondary Payload Adapter (VESPA) allows to embark in its internal position either two (2) micro satellites or one (1) larger mini satellite. The ASAP-S central position compartment and VESPA inner position compartment typically comprises a conical structure and a platform on which the small sat adaptor(s) is(are) attached. The configuration of these compartments is adjustable to the customer(s) need(s). Specific accommodations are available providing that the mass of the cluster of small sat remains in the qualification domain of the carrying structures.

Supplementary carrying systems will be also available for embarking Small Passengers on Vega-C and A6, in particular: • The SSMS (Small Satellite Mission Service) system for cluster launches on Vega and/or Vega-C, • The MLS (Microsat Launch Share) system for piggyback launches on A6.

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1.6.2. ASAP-S The ASAP-S system is the carrying structure allowing multiple launch configurations on Soyuz. It is mated on the Fregat upper stage, in lieu of the LVA-S, used in single launch configuration . The ASAP-S is manufactured by AIRBUS DS CASA. It consists of a load bearing carbon structure, comprising a cylindrical part and an upper truncated conical shell supporting the main passenger with its adaptor. On the side of the cylindrical part, four (4) external platforms are available.

I/F for main passenger adapter External platforms

Springs Inner platform Clamp Band

I/F with Fregat

Figure 1.6.2a ASAP-S configuration

Three versions of the ASAP-S are available: with 4 external platforms, with 2 external platforms only and with no external platforms. The separation of the upper part of the ASAP-S structure (with the main passenger adapter attached) is achieved by means of a Clamp Band and 8 springs.

Figure 1.6.2b ASAP-S – VS14 launch campaign

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1.6.3. VESPA The VEGA system is the carrying structure allowing multiple launch configurations on Vega. It is mated on the AVUM upper stage, in lieu of the standard 937 adaptor, used in single launch configuration . The VESPA is also manufactured by AIRBUS DS CASA. It consists of the upper part, the boat tail, the inner cone and the inner platform. It also provides the interface and separation system to the main passenger (Ø937 mm interface).

I/F and separation system for main passenger

Upper part Springs

Clamp Band Inner platform

Boat tail

I/F with AVUM

Figure 1.6.3a VESPA configuration Two versions of the VESPA are available: a standard version and a stretched version (height increased by 500mm). The separation of the upper part of the VESPA structure is achieved by means of a Clamp Band and 8 springs.

Figure 1.6.3b VESPA – VV07 launch campaign

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INTERFACES Chapter 2 FOR AUXILIARY PASSENGERS

2.1 General This chapter provides the Auxiliary Passenger interfaces with the adapter, carrying systems and on-board and ground electrical equipment. The spacecraft is mated to the launcher carrying system through an adapter that provides mechanical interface, electrical harnesses routing and systems to assure the spacecraft separation. Off-the-shelf adapters, with interface diameter of Ø298 mm, Ø381 mm, Ø432 mm, Ø610 mm and Ø937 mm are available for Micro and Mini Auxiliary Passengers. The electrical interface provides links with the launch vehicle and the ground support equipment during all phases of Auxiliary Passengers preparation. The payload fairing protects the spacecraft(s) from external environment during the flight as on the ground.

2.2 The reference axes All data and models shall be given in the same reference axis system to facilitate the interface configuration control and verification. The clocking of the spacecraft with regard to the adaptors, carrying structure and launch vehicle axes are defined by Arianespace taking into account main passenger requirements and Auxiliary Passengers characteristics (volume, electrical umbilical position, …). It is reported in the Interface Control Document.

2.3 Encapsulated spacecraft interfaces 2.3.1 General For auxiliary passenger, the available compartments are: - beneath the ASAP-S and VESPA structure, - on the external positions of the ASAP-S structure, between the ASAP-S cylindrical structure and the ST fairing.

2.3.2 Payload usable volume definition The payload usable volume is the area under the carrying structure or the fairing, available to the Auxiliary Passenger mated on its adapter. This volume constitutes the limits that the static dimensions of the spacecraft, including manufacturing tolerance and thermal protection, shall not exceed. One payload usable volume is defined for Mini Auxiliary Passengers and one payload usable volume is defined for Micro Auxiliary Passengers. They have been established having regard to the necessary margins at separation, to the potential displacement of the spacecraft during atmospheric phase and to the necessary accessibility to the mating interface during operations. They have been also established so as to ensure full compatibility between the ASAP-S and VESPA launch capability.

Arianespace ©, June 2017 §2 - p18 Auxiliary Passengers User’s Manual Interfaces for Auxiliary Passengers Issue 1 The above figures comprise the volume for spacecraft and its adaptor. The allocated volume envelope in the vicinity of the adapter is described in annex 2 for each of the off- the-shelf adapters.

2.3.2.2 Usable volume for Micro Auxiliary passenger The payload usable volume for Micro Auxiliary Passengers is shown below (for S/C and its adaptor):

Figure 2.3.2.2a: Micro Aux. Passenger Usable Volume (for S/C and its adaptor)

Figure 2.3.2.2b: Micro Aux. Passenger Figure 2.3.2.2c: Micro Aux. Usable Volume - ASAP-S on Soyuz Passenger Usable Volume - VESPA on VEGA

Note: The above usable volume is only valid with simultaneous separation of an even number of auxiliary passengers, similar in terms of mass.

For spacecraft with protrusions outside these dimensions, please contact Arianespace.

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Figure 2.3.2.3: Micro Aux. Passenger Extended Usable Volume on ASAP-S (for S/C and its adaptor)

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Figure 2.3.2.2b: Micro Aux. Passenger Extended Usable Volume - ASAP-S on Soyuz

Note1: The above usable volume is only valid on ASAP-S external positions and with simultaneous separation of an even number of auxiliary passengers, similar in terms of mass.

Note2: A circular opening with a diameter of 150 mm is provided in the external ASAP-S platforms for any protruding elements (boom, thruster, sensor, … for instance). Its location can be adapted at a certain extent.

2.3.2.3 Usable volume for Mini Auxiliary passenger The payload usable volume for Mini Auxiliary Passengers is shown below (for S/C and its adaptor):

Figure 2.3.2.3a: Mini Auxiliary Passenger Usable Volume (for S/C and its adaptor)

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Figure 2.3.2.3b: Mini Aux. Passenger Figure 2.3.3.1c: Mini Aux. Passenger Usable Volume - ASAP-S on Soyuz Usable Volume - stretched VESPA on VEGA

For spacecraft with protrusions outside these dimensions, please contact Arianespace.

2.3.3 On-fairing mission insignia One mission insignia based on Customer supplied artwork can be placed by Arianespace on the cylindrical section of the fairing. Standard insignia size is 600 x 600 mm square. The artwork shall be supplied not later than 6 months before launch.

2.4 Mechanical Interface Arianespace offers a range of standard off-the-shelf adapters compatible with most of the Auxiliary Passengers. All adapters are equipped with a payload separation system, brackets for electrical connectors. The electrical connector is mated on one bracket installed on the adapter and spacecraft side. On the spacecraft side, the umbilical connector bracket must be stiff enough to prevent any deformation greater than 0.5 mm under the maximum force of the connector spring. Depending on the adapter, it comprises the passive ring and the active ring, or the active ring only. The so-called passive ring is the I/F part remaining on the S/C. The so-called active ring is the I/F part remaining on the LV which comprises the active separation system. As a consequence, depending on the adapter, the S/C – LV interface can be the separation plan or a bolted I/F above the separation plan.

§2 - p22 Arianespace ©, June 2017 Auxiliary Passengers User’s Manual Interfaces for Auxiliary Passengers Issue 1 Standard adapters: The general characteristics of the off-the-shelf adapters are presented in the table below. A more detailed description is provided in Annex 2.

Adapter / Manufacturer Description Separation system SSASAP5 Dassault Active and passive ring S/C interface : 12*M6 bolts at Ø298 mm Total height: 91 mm Pyro-cutter ∅264 mm + up to 10 springs Total mass: 5 kg Mass retained on the S/C : 1 kg

PAS 381 S RUAG Space AB Active and passive ring S/C interface : 24 bolts at Ø381 mm Clamp-band ∅381 mm with Total height: 79 mm low shock separation system (CBOD-LT) + up to 24 Total mass: 4 kg springs Mass retained on the S/C : 1 kg

PAS 432 S RUAG Space AB Active and passive ring S/C interface : 12*M6 bolts at Ø298 mm Total height: 178 mm Total mass: 7 kg Clamp-band ∅432 mm with Mass retained on the S/C : 2 kg low shock separation system (CBOD-LT) + up to 6 springs

PAS 432 S RUAG Space AB Active ring alone S/C interface : Ø432 mm flange

Total height: 98 mm Total mass: 5 kg Mass retained on the S/C : 0 kg

PAS 610 S RUAG Space AB Clamp-band ∅610 mm with Active ring low shock separation system S/C interface : Ø610 mm flange CBOD-LT) + up to 12 springs Total height: 103 mm Total mass: 6 kg Mass retained on the S/C : 0 kg

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AR 937 AIRBUS DS Clamp -band ∅945 mm with low shock separation system Active ring S/C interface : Ø945 mm flange LPSS) + up to 8 springs Total height: 160 mm Total mass: 45 kg Mass retained on the S/C : 0 kg

Table 2.4.1 – Standard Adapters Note: In some situations, the Customer may wish to provide the payload adapter. In such cases, the Customer shall ask the Arianespace approval and corresponding requirements. Arianespace will supervise the design and production of such equipment to insure the compatibility at system level.

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2.5 Electrical interfaces The electrical links between the spacecraft, LV and the EGSE located at the preparation facilities and launch pad insure all needs of communication with Auxiliary Passengers during the launch preparation. Refer to Soyuz and Vega Users’ Manuals for the wiring diagram on the launch pad. The data and communication network used for spacecraft preparation in the CSG facilities are described in Chapter 6. The available electrical interface composition between spacecraft and the LV is presented in Table 2.5a below for Micro Auxiliary Passengers and in Table 2.5b for Mini Auxiliary Passengers.

Table 2.5a – Micro Auxiliary Passenger to launch vehicle electrical interfaces Provided Service Description Lines definition I/F connectors as 12 lines available Spacecraft TC/TM Umbilical per satellite data transmission Standard lines and battery charge see §2.5.1 1 × 12 pin DBAS 74 12 OSN LV Separation see §2.5.4 059 electrical monitoring functions Standard for spacecraft

Table 2.5b – Mini Auxiliary Passenger to launch vehicle electrical interfaces Provided Service Description Lines definition I/F connectors as 34 lines available Spacecraft TC/TM Umbilical per satellite data transmission Standard lines and battery charge see §2.5.2 2 × 37* pin

LV Separation see §2.5.4 DBAS 70 37 OSN electrical monitoring DBAS 70 37 OSY functions Standard for spacecraft

Note: * Smaller connectors (12, 19 or 27 pins) are also available on request.

Arianespace will supply the Customer with the spacecraft side interface connectors compatible with equipment of the off-the-shelf adaptors.

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2.5.1 Lines definition for Micro Auxiliary Passenger As a standard, for Micro Auxiliary passenger, 12 lines are available at the spacecraft- payload adapter interface via one umbilical link.

The standard 12 pins connector references are: - DBAS 74 12 OSN 059 on spacecraft side. - DBAS 025 82 10 12 on launcher side They are provided by Arianespace.

One pin is reserved for shielding.

Two pins are reserved for separation detection (separation status via the launch vehicle telemetry system).

The Customer has to define an electrical continuity using the separation system.

2.5.2 Lines definition for Mini Auxiliary Passenger As a standard, for Mini Auxiliary passenger, 34 lines are available at the spacecraft- payload adapter interface via two umbilical links.

The standard 37 pins connector references are: - DBAS 74 37 OSN 059 and DBAS 74 37 OSY 059 on spacecraft side. - DBAS 025 82 10 37 and DBAS 025 82 14 37 on launcher side. They are provided by Arianespace.

One pin per connector is reserved for shielding.

The Customer has to define an electrical continuity using the separation system.

2.5.3 Timeline limitation for Auxiliary Passengers Battery trickle charge is authorized up to J-2. The Auxiliary Passenger shall be inert (no RF emission, no status changes, …) during the final countdown and ascent phase up to a certain delay after the auxiliary passenger separation. The requested delay after separation will be determined by Arianespace depending on the type of orbit, mission timeline and main passenger requirements. Thus, no command can be sent to the auxiliary passenger, or generated by the auxiliary passenger onboard system (sequencer, computer, etc...).

2.5.4 Separation monitoring The separation status indication may be provided by dry loop straps integrated in the spacecraft/LV connector(s). The main electrical characteristics of these straps are: strap “closed”: R ≤ 1 Ω strap “open”: R ≥ 100 kΩ The dry loop straps can be: • dry loop straps located on adapter side, for the separation monitoring by satellite; • one dry loop strap located on satellite side, for the separation monitoring by the upper stage telemetry system.

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2.5.5 Electrical Continuity Interface

2.5.5.1 Bonding The spacecraft is required to have an “Earth” reference point close to the separation plane, on which a test socket can be mounted. The resistance between any metallic element of the spacecraft and a closest reference point on the structure shall be less than 10 mΩ for a current of 10 mA. The spacecraft structure in contact with the LV (separation plane of the spacecraft rear frame or mating surface of a Customer’s adapter) shall not have any treatment or protective process applied which creates a resistance greater than 10 mΩ for a current of 10 mA between spacecraft earth reference point and that of the LV (adapter or upper stage).

2.5.5.2 Shielding The umbilical shield links are grounded at both ends of the lines (the spacecraft on one side and EGSE on the other). The spacecraft umbilical grounding network diagram is shown in Soyuz and Vega Users’ Manuals. For each LV and ground harnesses connector, one pin is reserved to ensure continuity of the shielding.

2.6 Interface verifications 2.6.1 Prior to the launch campaign The Auxiliary Passenger can provide compatibility demonstration to mechanical and electrical interface either by heritage or with a mechanical and electrical fit check performed prior to the initiation of the launch campaign, as an option available at customer request. In case of a first application, the mechanical and electrical fit check prior to the launch campaign, is mandatory. Specific LV hardware for these tests is loaned according to the contractual provision.

The objectives of this fit-check are to confirm that the satellite dimensional and mating parameters meet all relevant requirements as well as to verify operational accessibility to the interface and cable routing. It can be followed by a release/drop test.

This test is usually performed at the Customer’s facilities, with flight representative launcher parts:

- the adapter equipped with its separation system and electrical connectors (provided by Arianespace)

S/C based on recurrent platform usually doesn’t need a fit-check prior to the launch campaign.

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2.6.2 Pre-launch validation of the electrical interface

2.6.2.1 Definition The electrical interface between satellite and launch vehicle is validated on each phase of the launch preparation where its configuration is changed or the harnesses are reconnected. These successive tests ensure the correct integration of the satellite with the launcher and help to pass the non reversible operations. There are up to three major configurations: • Spacecraft mated to its adapter; • Spacecraft with adapter mated to LV carrying structure; • Upper composite mated to launch vehicle. Depending on the test configuration, the flight hardware, the dedicated harness and/or the functional simulator will be used.

2.6.2.2 Spacecraft simulator The spacecraft simulator is used to simulate spacecraft functions during pre-integration tests: it shall be representative of the spacecraft output/input circuit that communicates with the adapter umbilical line. It will be provided by the Customer. It shall be integrated in a portable unit with a weight not higher than 25 kg and dimensions less than L600 × P600 × H800 mm. The simulator can be powered from external source.

2.6.2.3 Spacecraft EGSE The following Customer’s EGSE will be used for the interface validation tests: • OCOE, spacecraft test and monitoring equipment, permanently located in PPF Control rooms (LBC) and linked with the spacecraft during preparation phases and launch even at other preparation facilities and launch pad. • COTE, Specific front end Check-out Equipment, providing spacecraft monitoring and control, ground power supply and hazardous circuit’s activation. The COTE follows the spacecraft during preparation activity in PPF, HPF and Upper Composite Integration Facility. During launch pad operation the COTE is installed in the launch pad rooms. The spacecraft COTE is linked to the OCOE by data lines to allow remote control. • set of ground patch panel cables for satellite electrical umbilical lines verification. The installation interfaces as well as environmental characteristics for the COTE are described in the Chapter 6 of the Soyuz and Vega User’s Manuals.

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AUXILIARY PASSENGERS Chapter 3 LAUNCH MISSION

3.1 General The launch profile is driven by main passenger technical requirements. For Micro Auxiliary Passengers and Cubesats, the mission profile will be defined by Arianespace, in particular:

- The ascent profile (number and durations of the upper stage boost phases and of the coast phases, visibility from ground stations) will be defined by Arianespace.

- The auxiliary payload separation time will be defined by Arianespace. The auxiliary passenger shall be compatible with the worst case flight duration.

- The attitude pointing at separation will be defined by Arianespace. The Auxiliary Passenger has no means to impose a preferred orientation.

For Mini Auxiliary Passenger in central ASAP-S position or internal VESPA position: - The re-ignition capabilities of the Vega and Soyuz upper stages allow, at a certain extent depending on available performance margins, some modifications of the orbit altitude, eccentricity and inclination to attain the specific needs of the Mini Auxiliary Passenger. - The customer preferred separation conditions can be considered, depending on overall mission timeline.

3.2 Typical mission profile with auxiliary passenger(s) Following separation of the main passenger on its targeted orbit, the upper stage performs the necessary maneuvers to safely separate the auxiliary passengers.. The separation timeline will be defined by Arianespace based on main passenger requirements and Auxiliary Passengers mission. In any case, the Main Passenger is the first spacecraft separated. The overall mission duration of the mission depends on several parameters, including actual upper part configuration, the preferred orbit altitude, the availability of ground stations, etc… The auxiliary passenger shall be compatible with a separation time up to 7 hours after lift-off.

3.2.1 Mission timeline on Soyuz On Soyuz, several mission profiles can be envisaged: Short duration profile with a limited number of Fregat boosts: • Main Passenger separation → Micro Auxiliary Passengers separations → Fregat boost(s) → ASAP-S upper part jettisoning → Mini Auxiliary Passenger separation Long duration profile with many Fregat boosts:

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• Main Passenger separation → Fregat boost(s) → ASAP-S upper part jettisoning → Mini Auxiliary Passengers separations → Fregat boost(s) → Micro Auxiliary Passenger separation At the end of the mission, typically, a last Fregat boost triggers a controlled reentry of the last stage in the Earth’s atmosphere.

A typical short mission profile consists of the following phases: • Ascent phase to reach the main passenger orbit (including one or two first boosts of the upper stage) • Main passenger separation • Collision Avoidance maneuvers • Micro Auxiliary Passengers separations (2 by 2) • Collision Avoidance maneuvers • Fregat boost (one or two boosts) • ASAP-S upper part separation • Reorientation of the composite • Mini Auxiliary Passenger separation • Collision Avoidance maneuvers • Last Fregat boost (for a controlled reentry)

3.2.2 Mission timeline on Vega On Vega, a typical mission profile is the following: • Main Passenger separation → AVUM boost(s) → VESPA upper part jettisoning → Mini Auxiliary Passenger separation or Micro Auxiliary Passengers simultaneous separations After spacecraft separations, typically, a last AVUM boost triggers a controlled reentry of the last stage in the Earth’s atmosphere.

The mission profile consists of the following phases: • Ascent phase to reach the main passenger orbit (including one or two first boosts of the upper stage) • Main passenger separation • Collision Avoidance maneuvers • AVUM boost (one or two boosts) • VESPA upper part separation • Reorientation of the upper composite • Mini Auxiliary Passenger separation or, simultaneous Micro Auxiliary Passengers separations • Collision Avoidance maneuvers • Last AVUM boost (for a controlled reentry)

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3.3 Aerothermal flux at fairing jettisoning Jettisoning of the payload fairing can take place at different times depending on the main passenger aerothermal flux requirements. The Micro Auxiliary Passenger on the ASAP-S external positions shall be compatible with the main passenger requirement. Typically, fairing separation takes place when the aerothermal flux is lower than 1135 W/m 2 around 200 seconds from liftoff. Depending on the main passenger mission and launcher ascent profile, a second peak of flux may be encountered after fairing jettisoning. In such case, the maximum cumulated flux is 440 kJ/m 2.

3.4 Conditions at Auxiliary Passenger separation For auxiliary passenger, the separation conditions (time, direction, …) are defined by Arianespace. [NB: For mini Auxiliary Passenger in ASAP-S central position or VESPA internal position, and depending on overall mission, the customer preferred separation conditions can be considered.] With most of the small satellite adapters (refer to chapter 2.4 and annex 2 for a detailed description of the available adapters), the energy provided by each spring can be tuned, allowing counteracting the effect of spacecraft nominal static unbalance at spacecraft separation. For this purpose, the customer shall provide, at the time of the final mission analysis kick-off, its best estimation of the Center of Gravity location (nominal value and dispersions). Separation conditions (depointing, residual angular velocities after separation) is determined in the frame of the mission analysis. The auxiliary passenger shall be compatible with the sizing scenario in terms of depointing and residual angular velocity after separation.

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MICRO AUXILIARY PASSENGER DESIGN AND VERIFICATION Chapter 4 REQUIREMENTS

• - Spacecraft design and verification requirements

4.1 Introduction The design and dimensioning requirements that shall be taken into account by any Customer intending to launch a Micro Auxiliary Passenger compatible with the Arianespace Systems are detailed in this chapter.

4.2 Design Requirements for Micro Auxiliary Passenger 4.2.1 Safety Requirements The Customer is required to design the spacecraft in conformity with the CSG Safety Regulations, refer to the Payload Safety Handbook, CSG-NT-SBU-16687-CNES, Edition 1, Revision 1 dated 06 May 2015.

4.2.2 Selection of spacecraft materials The spacecraft materials must satisfy the following outgassing criteria: - Total Mass Loss (TML) ≤ 1 %; - Collected Volatile Condensable Material (CVCM) ≤ 0.1 %. measured in accordance with the procedure “ECSS-Q-70-02A”.

4.2.3 Micro S/C Mass Properties • The mass of a Micro Auxiliary Passenger must be in the range from 50 to 200 kg . • Center of gravity position:

o XG < 450 mm (from the mounting plane of the spacecraft) o The static unbalance of the spacecraft must stay within d ≤ 15 mm For satellites with characteristics outside these domains, please contact Arianespace. NOTE1: Large CoG offset can be counteracted by an adequate tuning of the separation system when off-the-shelf adapter is used. Refer to separation system description in Annex 2. NOTE2: The dispersion on the CoG offset shall be no more than 3 mm at the time of Final Mission Analysis Kick-off.

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4.2.4 Frequency Requirements for Micro Auxiliary Passenger To prevent any dynamic coupling with fundamental modes of the Launch Vehicle and Carrying System, the Micro Auxiliary Passenger shall be designed with a structural stiffness which ensures that the following requirements are fulfilled. In that case, the design limit loads given in next paragraph are applicable.

Lateral frequencies The fundamental frequency in the lateral axis of a spacecraft hard-mounted at the interface must be as follows: ≥ 45 Hz No local mode should be lower than the first fundamental frequencies. Longitudinal frequencies: The fundamental frequency in the longitudinal axis of a spacecraft hard-mounted at the interface must be as follows: ≥ 90 Hz No local mode should be lower than the first fundamental frequencies.

4.2.5 Design Loads During ground operations and flight, the spacecraft is subjected to various static and dynamic loads. The associated loads at spacecraft-to-adapter interface are defined by Quasi-Static Loads (QSL), that apply at spacecraft center of gravity and that are the most severe combinations of dynamic and static accelerations that can be encountered by the spacecraft at any instant of the mission. For a spacecraft complying with the stiffness requirements defined in previous paragraph 4.2.3, the limit levels of Quasi-Static Loads, to be taken into account for the design and dimensioning of the spacecraft primary structure, are given in Table 4.2.5 below:

QSL (g) (Static + Dynamic) Ground and Longitudinal Lateral flight load cases Compression Tension Case 1 - 14.2 + 10.3 +/- 2.6 Case 2 - 10.0 + 10.0 +/- 4.0

Table 4.2.5 - Design limit load for Micro Auxiliary Passenger Notes: • The factors apply on payload center of gravity, • The minus sign indicates compression along the longitudinal axis and the plus sign tension, • Lateral loads may act in any direction simultaneously with longitudinal loads, • The gravity load is included, • For the structural design, additional safety factors shall be applied as defined in paragraph 4.3.

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4.2.6 Line loads peaking on the spacecraft The geometrical discontinuities and differences in the local stiffness of the upper part structures (adapters and/or carrying structures) may produce local variations of the uniform line loads distribution. For the off-the-shelf adapters described in Annex 2, a value of 15% over the average line loads seen by the spacecraft is to be taken into account.

4.2.7 Line loads peaking induced by the Micro Auxiliary Passenger The maximum value of the peaking line load induced by the spacecraft is allowed in local areas to be up to 50% over the maximum line loads induced by the dimensioning loads (deduced from QSL table 4.2.5.). 4.2.8 Handling loads during ground operations During the encapsulation phase, the spacecraft is lifted and handled with its adapter: for this reason, the spacecraft and its handling equipment must be capable of supporting an additional mass of 15 kg .

4.2.9 Local loads On top of the global loads described in the above paragraphs, local loads shall be considered for spacecraft sizing, including payload adapter separation spring forces and flatness effect at spacecraft-to-adapter interface.

4.2.10 Dynamic loads The secondary structures and flexible elements (e.g., solar panels, antennas, and propellant tanks) must be designed to withstand the dynamic environment with the appropriate safety factors as defined in paragraph 4.3.

4.3 Micro Satellites compatibility verification requirements

During the preparation for launch and during the flight, the spacecraft is exposed to a variety of mechanical, thermal and electromagnetic environments. Refer to Soyuz user’s Manual and Vega User’s Manual chapter 3 for an extensive description of these environments. The present Chapter describes the requested demonstrations applicable for Micro Auxiliary Passenger.

4.3.1 Verification Logic The spacecraft authority shall demonstrate that the spacecraft structure and equipments are capable of withstanding the maximum expected launch vehicle ground and flight environments. The spacecraft compatibility must be proven by means of adequate tests. The verification logic with respect to the satellite development program approach is shown in Table 4.3.1 below:

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S/C development Model Static Sine Random Acoustic Shock approach vibration vibration

Shock test STM Qual. test Qual. test Qual. test N/A (4) characteriz ation and

analysis

With Structural Shock test (4) Test Model FM1 By heritage Protoflight Protoflight N/A characteriz (STM) from STM (1) test (2) Test (2) ation and analysis or by heritage (1)

By heritage Acceptance Acceptance Subsequent from STM (1) test test N/A (4) By FM’s (3) (optional) (optional) heritage and analysis (1)

Qual. test or Protoflight Protoflight N/A (4) Shock test With PFM = FM1 by heritage test (2) Test (2) characteriz ProtoFlight (1) ation and Model analysis or (PFM) by heritage (1)

Acceptance Acceptance N/A (4) By Subsequent By heritage test test heritage FM’s (3) (1) (optional) (optional) and analysis (1) Table 4.3.1 – Spacecraft verification logic

Notes: (1) If qualification is claimed by heritage, the representativeness of the structural test model (STM) with respect to the actual flight unit must be demonstrated. (2) Protoflight approach means qualification levels and acceptance duration/sweep rate. (3) Subsequent FM: spacecraft identical to FM1 (same primary structure, major subsystems and appendages). (4) Acoustic test not required for micro satellites (refer to chapter 4.3.4.5) pending demonstration that the acoustic environment is covered by the random test.

The mechanical environmental test plan for spacecraft qualification and acceptance shall comply with the requirements presented hereafter and shall be reviewed by Arianespace prior to implementation of the first test.

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The purpose of ground testing is to screen out unnoticed design flaws and/or inadvertent manufacturing and integration defects or anomalies. It is therefore important that the satellite be mechanically tested in flight-like configuration. In addition, should significant changes affect the tested specimen during subsequent AIT phase prior to spacecraft shipment to CSG, the need to re-perform some mechanical tests must be reassessed. If, despite of notable changes, complementary mechanical testing is not considered necessary by the Customer, this situation should be treated in the frame of a Request For Waiver, which justification shall demonstrate, in particular, the absence of risk for the launcher.

4.3.2 Safety factors Spacecraft qualification and acceptance test levels are determined by increasing the limit loads by the safety factors given in Table 4.3.2 below. The spacecraft must have positive margins with these safety factors.

Qualification (4) Protoflight Acceptance

SC tests Factors Duration/ Factors Duration/ Factors Duration/ Rate Rate Rate

Static 1.25 N/A 1.25 N/A N/A N/A (QSL)

0.5 1.0 1.0 Sine 1.25 oct./min 1.25 oct./min 1.0 oct./min vibrations (2) (2) (2)

Random 2.25 (1) 240 s 2.25 (1) 120 s 1.0 (1) 120 s vibrations

+3 dB +3 dB Acoustics 120 s 60 s 1.0 60 s (or 2) (or 2)

+3 dB N/A +3 dB N/A Shock N/A (or 1.41) (3) (or 1.41) (3)

Table 4.3.2 - Test Factors, rate and duration

Notes: (1) Factor by which to multiply the Power Spectral Density. (2) See paragraph 4.3.3.2. (3) Number of tests to be defined in accordance with methodology for qualification (see paragraph 4.3.3.5.). (4) If qualification is not demonstrated by test, it is reminded that a safety factor of 2 is requested with respect to the design limit.

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4.3.3. Spacecraft compatibility tests for micro auxiliary passenger 4.3.3.1. Static tests Static load tests (in the case of an STM approach) are performed by the Customer to confirm the design integrity of the primary structural elements of the spacecraft platform. Test loads are based on worst-case conditions, i.e. on events that induce the maximum mechanical line loads into the main structure, derived from the table of maximum QSLs (paragraph 4.2.5) and taking into account the additional line loads peaking (paragraph 4.2.6) and the local loads (paragraph 4.2.9). The qualification factors (paragraph 4.3.2) shall be considered.

4.3.3.2. Sinusoidal vibration tests The objective of the sine vibration tests is to verify the spacecraft secondary structure dimensioning under the flight limit loads multiplied by the appropriate safety factors. The spacecraft qualification test consists of one sweep through the specified frequency range and along each axis. The levels to be applied are presented in Table 4.3.5.2 and illustrated in Figure 4.3.5.2 below.

Frequency Qualification Acceptance Sine range (Hz) levels (0-peak) g levels (0-peak) g 5 – 75 2.50 2.0 Longitudinal 75 – 110 1.25 1.0 110 - 125 0.25 0.2 5 – 60 1.87 1.5 Lateral 60 – 110 0.62 0.5 110 - 125 0.25 0.2

Table 4.3.3.2 – Sinusoidal vibration tests levels (micro S/C)

Figure 4.3.3.2 – Sinusoidal vibration tests acceptance levels (micro S/C)

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A notching procedure may be agreed in the frame of a request for waiver, on the basis of the latest coupled loads analysis (CLA) available at the time of the tests to prevent excessive loading of the spacecraft structure. However it must not jeopardize the tests objective to demonstrate positive margins of safety with respect to the flight limit loads, while considering appropriate safety factor. In addition a sweep rate increase may be agreed in the frame of a request for waiver to limit fatigue solicitations during the test. The acceptability of the sweep rate shall consider the dynamic characteristics of spacecraft secondary structures or appendages and the actual damping of the payload structure, in order to ensure proper solicitation of the whole spacecraft during the test.

4.3.3.3. Random vibration tests The verification of the spacecraft structure compliance with the random vibration environment in the 20Hz - 2000Hz frequency range shall be performed. The levels to be applied are presented in Table 4.3.5.3 and illustrated in Figure 4.3.5.3 below.

Spectral density (10 -3 g2/Hz) Frequency band Qualification Acceptance 20 – 50 11.25 5 50 – 100 11.25 – 22.5 5 – 10 100 – 200 22.5 – 56.25 10 – 25 200 – 500 56.25 25 500 – 1000 56.25 – 22.5 25 – 10 1000 – 2000 22.5 – 11.25 10 – 5 Overall (g) 7.5 5

Table 4.3.3.3 – Random vibration tests levels (micro S/C)

Figure 4.3.3.3 – Random vibration tests acceptance levels (micro S/C)

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4.3.3.4. Acoustic vibration tests For Micro Auxiliary Passenger, the acoustic environment is covered by the random test. Refer to Chapter 5 Paragraph 5.3.5.4. for further details on the acoustic environment.

4.3.3.5. Shock qualification The spacecraft is subject to shock primarily during stage separations, Fairing jettisoning and actual spacecraft separation.

The envelope acceleration shock response spectrum (SRS) at the spacecraft base (computed with a Q-factor of 10) is presented in Tables 4.3.3.5.1 to & 4.3.3.5.3 and Figure 4.3.3.5.1. These levels are applied simultaneously in axial and radial directions.

Flight Event Frequency (Hz) 100 – 1600 1600 – 10000 SRS, Shock Response Spectra (Q = 10) (g) Fairing & stages 30 – 2000 2000 separations Table 4.3.3.5.1 - Shock response spectra at Fairing and stages separations Flight Event Frequency (Hz) 100 – 1000 1000 – 10000 Clamp band spacecraft SRS, Shock Response Spectra (Q = 10) (g) separation Ø381, Ø432, 20 – 1000 1000 – 700 Ø610 Table 4.3.3.5.2 - Shock response spectra for clamp band separation (PAS) Frequency (Hz) Flight Event 200–2000 2000– 8000–

8000 10000 SRS, Shock Response Spectra (Q = 10) (g) Pyro spacecraft separation 60 – 3500 3500 – 4500 4500 Table 4.3.3.5.3 - Shock response spectra for pyro separation (SSASAP5)

Figure 4.3.3.5 – Envelope shock response spectra (SRS) at the spacecraft base

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The ability of the spacecraft to withstand the shock environment generated by the stages separation, the fairing jettisoning and the spacecraft separation shall follow a comprehensive process including tests and analysis.

¶ Clamp-band separation event

When an adapter with a clampband is used, the demonstration of the spacecraft’s ability to withstand the separation shock generated by the clamp-band release shall be based on one of the following methods:

Method Number One: Release drop test, extrapolation to specification, comparison to S/C sub-systems qualification:

A clamp-band release drop test is conducted with the tension of the band set at the nominal tension at installation. During this test, interface levels and equipment base levels are measured. This test can be performed on the STM, on the PFM or on the first flight model provided that the spacecraft structure close to the interface as well as the equipment locations and associated supports are equivalent to those of the flight model. The release shocks generated at the spacecraft’s interface and measured during the above-mentioned test are compared to the applicable shock specification (see above, Table 4.3.3.5.2). The ratio derived from the above comparison is then considered to extrapolate the measured equipment levels to the specification. These extrapolated shock levels are then increased by a safety factor of +3 dB and are compared to the qualification status of each spacecraft subsystem and/or equipment. Note that each unit qualification status can be obtained from environmental qualification tests other than shock tests by using equivalent rules (e.g. from sine or random vibration tests).

Method Number Two: Release drop test with maximal tension, direct comparison to S/C sub-systems qualification:

A clamp-band release drop test is conducted with the tension of the band set as close as possible to its maximum value during flight. During this test, interface levels and equipment base levels are measured. This test can be performed on the STM, on the PFM or on the first flight model provided that the spacecraft structure close to the interface as well as the equipment locations and associated supports are equivalent to those of the flight model. The induced shocks generated on spacecraft equipment measured during the above mentioned test are then increased by: - A +3 dB uncertainty margin aiming at deriving flight limit environment from the single test performed in flight-like configuration; [NB: In case two clamp-band release drop tests are performed, this +3 dB uncertainty margin can be removed but the maximum recorded value between the two tests has to be considered for each equipment.] - A +3 dB safety factor aiming at defining the required minimum qualification levels, to be compared to the qualification status of each spacecraft subsystem and/or equipment.

These obtained shock levels are then compared to the qualification status of each spacecraft subsystem and/or equipment. Note that each unit qualification status can be obtained from environmental qualification tests other than shock tests by using equivalent rules (e.g. from sine or random vibration tests).

¶ Pyro separation event

When a SSASAP5 adapter is used, the demonstration of the spacecraft’s ability to withstand the separation shock generated by the pyro release shall be based on the following method:

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Method: shock test using shock test device, direct comparison to S/C sub-systems qualification:

A shock test device can be provided by Arianespace, to simulate the pyro shock. It is composed of four shock generators actuated simultaneously and an interface shock plate to distribute the shock wave on the satellite. The number of shock generators provided allows to perform three tests. During these tests, interface levels and equipment base levels are measured. These tests can be performed on the STM, on the PFM or on the first flight model provided that the spacecraft structure close to the interface as well as the equipment locations and associated supports are equivalent to those of the flight model. The induced shocks generated on spacecraft equipment measured during the above mentioned test* are then increased by a +3 dB safety factor aiming at defining the required minimum qualification levels. These obtained shock levels are then compared to the qualification status of each spacecraft subsystem and/or equipment. Note that each unit qualification status can be obtained from environmental qualification tests other than shock tests by using equivalent rules (e.g. from sine or random vibration tests).

* The maximum recorded value between the three tests has to be considered for each equipment.

¶ Launcher events (fairing/stages separation)

For a Vega launch, an additional shock test shall be performed in order to characterize the shock transmission inside the spacecraft and define the transfer functions between the spacecraft interface plane and the equipment base, when the S/C remains attached to the launcher.

As above, this test can be performed on the STM, PFM or on the first flight model, provided that the spacecraft configuration is representative of the flight model (structure, load paths, equipment presence and location,…). This test can be performed once, and the verification performed covers the spacecraft platform as far as no structural modification alters the validity of the analysis. This qualification is obtained by comparing the component unit qualification levels to the equipment base levels experienced applying the interface shock spectrum specified above, Table 4.3.3.5.1 with the dedicated transfer function. A minimum +3 dB margin has to be highlighted to validate the qualification (see Table 4.3.1). Note that each unit qualification status can be obtained from environmental qualification tests other than shock tests by using equivalent rules (e.g. from sine or random vibration tests).

For a Vega launch, the VEga Shock Test Apparatus (VESTA) generating a shock environment representative of the actual Vega fairing separation event can be provided by Arianespace. Thanks to the representativeness of this test mean, the spacecraft qualification can be directly derived from the VESTA tests results, removing the uncertainties margins taken into consideration.

General nota: In case of recurring platform or spacecraft, the shock qualification can be based on heritage, pending that identical platform or spacecraft is already qualified to both launcher and clamp-band release event.

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4.4. Thermal loads The thermal environment during ground operations and during ascent phase is described in Soyuz and Vega User’s Manuals (refer to corresponding chapters 3.4). During the ground phase, the fairing ventilation conditions (air flow rate and air temperature setting point) are set to fulfill main passenger requirements.

During the ascent phase, the time of auxiliary passenger separation can be up to lift-off + 7 Hours. It is determined by Arianespace taking into account the optimization being made for the Main Passenger and for the overall mission (e.g. separation of other auxiliary passengers and de-orbiting at the end of mission). The S/C shall sustain the associated thermal conditions.

4.5. RF environment The Micro Auxiliary Passenger emitters shall be OFF during all the Combined operations, final countdown and ascent phase up to a certain delay after the auxiliary passenger separation. The requested delay after separation will be determined by Arianespace depending on the type of orbit, mission timeline and main passenger requirements.

The intensity of the electrical field generated by spurious or intentional emissions from the launch vehicle and the range RF systems does not exceed the following levels:

Figure 4.5 – Spurious radiation by LV and launch base narrow-band electrical field

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MINI AUXILIARY PASSENGER DESIGN AND VERIFICATION Chapter 5 REQUIREMENTS

5.1 Introduction The design and dimensioning requirements that shall be taken into account by any Customer intending to launch a Mini Auxiliary Passenger compatible with the Arianespace Systems are detailed in this chapter.

5.2 Design requirements for Mini Satellites 5.2.1 Safety Requirements The Customer is required to design the spacecraft in conformity with the CSG Safety Regulations, refer to the Payload Safety Handbook, CSG-NT-SBU-16687-CNES, Edition 1, Revision 1 dated 06 May 2015.

5.2.2 Selection of spacecraft materials The spacecraft materials must satisfy the following outgassing criteria: - Total Mass Loss (TML) ≤ 1 %; - Collected Volatile Condensable Material (CVCM) ≤ 0.1 %. measured in accordance with the procedure “ECSS-Q-70-02A”.

5.2.3 Mini S/C Mass Properties • The mass of a Micro Auxiliary Passenger must be in the range from 200 to 400 kg . • Center of gravity position:

o XG < 900 mm (from the mounting plane of the spacecraft) o The static unbalance of the spacecraft must stay within d ≤ 15 mm For satellites with characteristics outside these domains, please contact Arianespace. NOTE1: Large CoG offset can be counteracted by an adequate tuning of the separation system when off-the-shelf adapter is used. Refer to separation system description in Annex 2. NOTE2: The dispersion on the CoG offset shall be no more than 5 mm at the time of Final Mission Analysis Kick-off.

5.2.3 Frequency Requirements for Mini Auxiliary Passenger To prevent dynamic coupling with fundamental modes of Launch Vehicle and Carrying System, the Mini Auxiliary Passenger should be designed with a structural stiffness which ensures that the following requirements are fulfilled. In that case, the design limit loads given in next paragraph are applicable.

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Lateral frequencies The fundamental frequency in the lateral axis of a spacecraft hard-mounted at the interface must be as follows: ≥ 20 Hz No local mode should be lower than the first fundamental frequencies. Longitudinal frequencies: The fundamental frequency in the longitudinal axis of a spacecraft hard-mounted at the interface must be as follows: ≥ 60 Hz No local mode should be lower than the first fundamental frequencies.

5.2.4 Design Loads During ground operations and flight, the spacecraft is subjected to various static and dynamic loads. The associated loads at spacecraft-to-adapter interface are defined by Quasi-Static Loads (QSL), that apply at spacecraft center of gravity and that are the most severe combinations of dynamic and static accelerations that can be encountered by the spacecraft at any instant of the mission. For a spacecraft complying with the stiffness requirements defined in previous paragraph 5.2.3, the limit levels of Quasi-Static Loads, to be taken into account for the design and dimensioning of the spacecraft primary structure, are given in Table 5.2.5 below:

QSL (g) (Static + Dynamic) Ground and Longitudinal Lateral flight load cases Compression Tension - 7.5 + 5.0 +/- 2.6

Table 5.2.5 - Design limit load for Mini Auxiliary Passenger

Notes: • The factors apply on payload center of gravity, • The minus sign indicates compression along the longitudinal axis and the plus sign tension, • Lateral loads may act in any direction simultaneously with longitudinal loads, • The gravity load is included, • For the structural design, additional safety factors shall be applied as defined in paragraph 5.3.

5.2.5 Line loads peaking on the spacecraft The geometrical discontinuities and differences in the local stiffness of the upper part structures (adapters and/or carrying structures) may produce local variations of the uniform line loads distribution. For the off-the-shelf adapters described in Annex 2, a value of 15% over the average line loads seen by the spacecraft is to be taken into account.

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5.2.6 Line loads peaking induced by the Mini Auxiliary Passenger The maximum value of the peaking line load induced by the spacecraft is allowed in local areas to be up to 20% over the maximum line loads induced by the dimensioning loads (deduced from QSL Table 5.2.5.). 5.2.7 Handling loads during ground operations During the encapsulation phase, the spacecraft is lifted and handled with its adapter: for this reason, the spacecraft and its handling equipment must be capable of supporting an additional mass of ~20 kg (Ø610 I/F) or additional mass of ~60 kg (Ø937 I/F).

5.2.8 Local loads On top of the global loads described in the above paragraphs, local loads shall be considered for spacecraft sizing, including payload adapter separation spring forces and flatness effect at spacecraft-to-adapter interface.

5.2.9 Dynamic loads The secondary structures and flexible elements (e.g., solar panels, antennas, and propellant tanks) must be designed to withstand the dynamic environment with the appropriate safety factors as defined in paragraph 5.3.

5.3 Spacecraft compatibility verification requirements During the preparation for launch and during the flight, the spacecraft is exposed to a variety of mechanical, thermal and electromagnetic environments. Refer to Soyuz user’s Manual and Vega User’s Manual chapter 3 for an extensive description of these environments. The present Chapter describes the requested demonstrations applicable for Micro Auxiliary Passenger.

5.3.1 Verification Logic The spacecraft authority shall demonstrate that the spacecraft structure and equipments are capable of withstanding the maximum expected launch vehicle ground and flight environments. The spacecraft compatibility must be proven by means of adequate tests. The verification logic with respect to the satellite development program approach is shown in Table 5.3.1 below:

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S/C development Model Static Sine Random Acoustic Shock approach vibration vibration

Shock test STM Qual. test Qual. test Qual. test Qual. test characteriz ation and

analysis

With Structural Shock test Test Model FM1 By heritage Protoflight Protoflight Protoflight characteriz (STM) from STM (1) test (2) Test (2) test (2) ation and analysis or by heritage (1)

By heritage Acceptance Acceptance Subsequent from STM (1) test test Acceptance By FM’s (3) (optional) (optional) test heritage and analysis (1)

Qual. test or Protoflight Protoflight Protoflight Shock test With PFM = FM1 by heritage test (2) Test (2) test (2) characteriz ProtoFlight (1) ation and Model analysis or (PFM) by heritage (1)

Acceptance Acceptance Acceptance By Subsequent By heritage test test test heritage FM’s (3) (1) (optional) (optional) and analysis (1) Table 5.3.1 – Spacecraft verification logic

Notes: (1) If qualification is claimed by heritage, the representativeness of the structural test model (STM) with respect to the actual flight unit must be demonstrated. (2) Protoflight approach means qualification levels and acceptance duration/sweep rate. (3) Subsequent FM: spacecraft identical to FM1 (same primary structure, major subsystems and appendages). The mechanical environmental test plan for spacecraft qualification and acceptance shall comply with the requirements presented hereafter and shall be reviewed by Arianespace prior to implementation of the first test.

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The purpose of ground testing is to screen out unnoticed design flaws and/or inadvertent manufacturing and integration defects or anomalies. It is therefore important that the satellite be mechanically tested in flight-like configuration. In addition, should significant changes affect the tested specimen during subsequent AIT phase prior to spacecraft shipment to CSG, the need to re-perform some mechanical tests must be reassessed. If, despite of notable changes, complementary mechanical testing is not considered necessary by the Customer, this situation should be treated in the frame of a Request For Waiver, which justification shall demonstrate, in particular, the absence of risk for the launcher.

5.3.2 Safety factors Spacecraft qualification and acceptance test levels are determined by increasing the limit loads by the safety factors given in Table 5.3.2 below. The spacecraft must have positive margins with these safety factors.

Qualification (4) Protoflight Acceptance

SC tests Factors Duration/ Factors Duration/ Factors Duration/ Rate Rate Rate

Static 1.25 N/A 1.25 N/A N/A N/A (QSL)

0.5 1.0 1.0 Sine 1.25 oct./min 1.25 oct./min 1.0 oct./min vibrations (2) (2) (2)

Random 2.25 (1) 240 s 2.25 (1) 120 s 1.0 (1) 120 s vibrations

+3 dB +3 dB Acoustics 120 s 60 s 1.0 60 s (or 2) (or 2)

+3 dB N/A +3 dB N/A Shock N/A (or 1.41) (3) (or 1.41) (3)

Table 5.3.2 - Test Factors, rate and duration

Notes: (1) Factor by which to multiply the Power Spectral Density. (2) See paragraph 4.3.3.2. (3) Number of tests to be defined in accordance with methodology for qualification (see paragraph 4.3.3.5.). (4) If qualification is not demonstrated by test, it is reminded that a safety factor of 2 is requested with respect to the design limit.

5.3.3 Spacecraft compatibility tests for mini auxiliary passenger 5.3.3.1 Static tests

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Static load tests (in the case of an STM approach) are performed by the Customer to confirm the design integrity of the primary structural elements of the spacecraft platform. Test loads are based on worst-case conditions, i.e. on events that induce the maximum mechanical line loads into the main structure, derived from the table of maximum QSLs (paragraph 5.2.5) and taking into account the additional line loads peaking (paragraph 5.2.6) and the local loads (paragraph 5.2.9). The qualification factors (paragraph 4.3.2) shall be considered.

5.3.3.2 Sinusoidal vibration tests The objective of the sine vibration tests is to verify the spacecraft secondary structure dimensioning under the flight limit loads multiplied by the appropriate safety factors. The spacecraft qualification test consists of one sweep through the specified frequency range and along each axis. The qualification levels to be applied are presented in Table 5.3.3.2 below.

Frequency Qualification Acceptance Sine range (Hz) levels (0-peak) g levels (0-peak) g 5 – 25 1.0 0.8 25 - 45 1.0 0.8 Longitudinal 45 - 60 2.5 2.0 60 - 110 1.25 1.0 110 - 125 0.25 0.2 5 - 25 1.0 0.8 25 - 45 0.62 0.5 Lateral 45 - 60 1.87 1.5 60 - 110 0.62 0.5 110 - 125 0.25 0.2

Table 5.3.3.2 – Sinusoidal vibration tests levels (Mini S/C)

Figure 5.3.3.2 – Sinusoidal vibration tests levels (Mini S/C)

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A notching procedure may be agreed in the frame of a request for waiver, on the basis of the latest coupled loads analysis (CLA) available at the time of the tests to prevent excessive loading of the spacecraft structure. However it must not jeopardize the tests objective to demonstrate positive margins of safety with respect to the flight limit loads, while considering appropriate safety factor. In addition a sweep rate increase may be agreed in the frame of a request for waiver to limit fatigue solicitations during the test. The acceptability of the sweep rate shall consider the dynamic characteristics of spacecraft secondary structures or appendages and the actual damping of the payload structure, in order to ensure proper solicitation of the whole spacecraft during the test.

5.3.3.3 Random vibration qualification The verification of the spacecraft structure compliance with the random vibration environment in the 20Hz - 2000Hz frequency range shall be performed. The levels to be considered are presented in Table 5.3.3.3 and illustrated in Figure 5.3.3.3 below.

Spectral density (10 -3 g2/Hz) Frequency band Qualification Acceptance 20 – 50 11.25 5 50 – 100 11.25 – 22.5 5 – 10 100 – 200 22.5 – 56.25 10 – 25 200 – 500 56.25 25 500 – 1000 56.25 – 22.5 25 – 10 1000 – 2000 22.5 – 11.25 10 – 5 Overall (g) 7.5 5

Table 5.3.3.3 – Random vibration levels (mini S/C)

Figure 5.3.3.3 – Random vibration acceptance levels (mini S/C)

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Two methodologies can be contemplated:

Method Number One: Perform a dedicated random vibration qualification test with the qualification levels in Table 5.3.3.3 above.

Method Number Two: Based on the spacecraft structural transfer functions derived from sine vibrations tests up to 2000Hz (with low levels from 125 to 2000 Hz), demonstrate by analysis the compliance of the spacecraft secondary structure with the random vibration environment in the [20-2000]Hz frequency range. The different steps to follow are: a) Restitution of the transfer functions TFi(f) between S/C interface and each equipment “i”:

where: TFi(f) – transfer function at location “i” Ri(f) – measured response at location “i” from sine test RI/F(f) – average of the pilot at S/C interface

b) Calculation of the random responses for each equipment “i”:

where: PSDi(f) – Power Spectral Density at location “i”, [g2/Hz] PSDinput(f) – input Power Spectral Density (accept. level) at f, [g2/Гц] QF – Qualification Factor = 1.5

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5.3.3.4 Acoustic vibration tests

Acoustic testing is accomplished in a reverberant chamber. The volume of the chamber with respect to that of the spacecraft shall be sufficient so that the applied acoustic field is diffuse. The test measurements shall be performed at a minimum distance of 1 m from spacecraft. Flight Limit Level (dB) -5 Octave Center (reference: 0 dB= 2 x 10 Pa) Test Frequency tolerance * (Hz) Qualification Acceptance 31.5 133 130 -2, +4 63 138 135 -1, +3 125 137 134 -1, +3 250 141.5 138.5 -1, +3 500 138.5 135.5 -1, +3 1000 130 127 -1, +3 2000 124 121 -1, +3 OASPL (20-2828 Hz) 145.5 142.5 Test duration 120s 60s

Table 5.3.3.4 – Acoustic vibration test levels

* Dispersion allowed between the microphone measurements. For each octave, the levels to be achieved in the anechoic chamber are applicable to the average of the microphones’ measurements.

No fill factor correction is applied.

Figure 5.3.3.4 – Acoustic vibration test levels

5.3.3.5 Shock qualification The shock environment and the logic for demonstrating spacecraft qualification to the shock environment are identical for micro and mini satellites. Refer to Chapter 4 Paragraph 4.3.3.5.

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5.4 Thermal loads The thermal environment during ground operations and during ascent phase is identical for micro and mini satellites.

Refer to Chapter 4 Paragraph 4.4

5.5 RF environment The intensity of the electrical field generated by spurious or intentional emissions from the launch vehicle and the range RF systems does not exceed the following levels:

Figure 5.5 – Spurious radiation by LV and launch base narrow-band electrical field

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MISSION MANAGEMENT & LAUNCH CAMPAIGN ORGANISATION FOR AUXILIARY PASSENGERS Chapter 6

6.1 Introduction The overall launch service for auxiliary passengers (management, mission integration process, hardware supply, launch campaign organization, range support) is briefly described in the present chapter.

6.2 Mission management 6.2.1 Contract organization The contractual commitments between the launch service provider and the Customer are defined in the Launch Services Agreement (LSA) with its Statement of Work (SOW), and its Technical Specification. Based on the Application to Use Arianespace launch vehicles (“DUA “ – refer to annex 1 for a template) filled out by the Customer, the Statement of Work identifies the task and deliveries of the parties, and its Technical Specification identifies the technical interfaces and requirements. The Arianespace Program Director appointed to auxiliary passenger(s) will provide the organization and resources to fulfill the contractual obligations: contract amendments, payments, planning, configuration control, documentation, reviews, meetings, etc… Additionally, during the launch campaign, the appointed Mission Director will handle the operations activities of the auxiliary passenger(s). 6.2.2 Schedule The schedule for auxiliary passenger will be established in compliance with the main passenger requirements and with the milestones specified in the Statement of Work of the Launch Service Agreement. In practice, the final activities (from L-10) shall follow the main passenger schedule (data delivery for final analysis, RAMF, BT-POC, etc…). 6.2.3 Meetings and Reviews A typical content of the different meetings and reviews is given below.

# Title Typical Object ç Loc. é date å 1 Contractual Kick-off Meeting: L - 24 M-E C Project management – project milestones – organisation – security & confidentiality aspects – communications protocol 2 DUA Review: L - 22 M-E-O-S E or W Review of the Spacecraft characteristics and requirements Review of the DCI Issue 0 Revision 0 3 Option : Prelim. Mission Analysis Review L - 17 M-E-O-S E (RAMP): As necessary, Orbit and injection accuracy –

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Separation and collision avoidance – Dynamic environment Review of Safety Submission status DCI review 4 DCI Signature Issue 1 Revision 0 L - 16 M-E-O-S T 5 Option: CSG site survey / Preparation of L - 14 M-O-S W or K S/C Operations Plan (POS): Launch base facilities visit – CSG support – Telecommunications network - Safety submission phases 1 and 2 DCI review (chapters 7 and 8) 6 Final Mission Analysis Kick-Off: L - 10 M-E T Review of the Final Mission Analysis inputs DCI review 7 Review of S/C Operations Plan (POS): L - 6 M-O-S W Transport and logistics – Preliminary S/C Operations Plan (POS) – Combined Operations introduction – Telecommunications network Safety submission phases 1 and 2 DCI review (chapters 7 and 8) 8 Final Mission Analysis Review (RAMF): L - 3 M-E E Trajectory – performance – injection accuracy – separation and collision avoidance – thermal – dynamic environment – EMC environment – authorization to start the flight program production – Spacecraft qualification status DCI review 9 Final Campaign Preparation Meeting: L - 3 M-O-S E Campaign preparation status - S/C Operations Plan (POS) – Interleaved Operations Plan (POI) – Combined Operations Plan Safety submission status DCI review (chapters 7 and 8) 10 DCI Signature Issue 2 Revision 0 L - 2 M-E-O-S T 11 POC Readiness Review (BT POC): è M-O-S K Launch Vehicle and Launch System status Spacecraft status 12 Financial Wrap-up meeting L-1 day M K å Dates are given in months, relative to L, where L is the first day of the Launch Term, Period, Slot or Day. ç M Ù Management; E Ù Engineering; O Ù Operations; S Ù Safety é E Ù Evry; K Ù Kourou; C Ù Customer Headquarter; W Ù S/C Manufacturer Plant; T Ù By Teleconference/Electronic exchange è To be held the day before the agreed day for starting the Combined Operations

The auxiliary passenger is also invited to attend the launcher reviews, RAV (LV Flight Readiness Review) and RAL (Launch Readiness Review). The L/V-S/C interfaces will be examined with reference to the DCI.

At the LV Flight Readiness Review, the Customer is asked to present the proof of the availability of the satellite Dummy.

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6.3 Systems engineering support The Arianespace’s launch service for auxiliary passenger includes the standard engineering tasks conducted to ensure Auxiliary Passenger compatibility with the main passenger mission and with the launch system. 6.3.1 Interface Management The technical interface management is based on the Interface Control Document (DCI “Document de Contrôle d’Interface”), which is prepared by Arianespace using inputs from the Technical Specification of the Launch Service Agreement.

6.3.2 Mission Analysis Mission analysis is typically organized into two phases, which are: • the Preliminary Mission Analysis, as option, to be exercised at the DUA review at the latest; and • the Final Mission Analysis. The Final Mission Analysis focuses on the actual flight configuration and uses the final actual data and models of the main and auxiliary passengers. The main outcomes of the analyses are collected in the Interface Control Document dedicated to the auxiliary passenger.

6.3.3 Auxiliary Passenger Compatibility Verification Compatibility to launch system environment: In close relationship with mission analysis, Arianespace will check the Auxiliary Passenger design is able to withstand the LV environment. For this purpose, the following reports will be required for review and approval: • An Auxiliary Passenger environment test plan correlated with requirements described in previous Chapters. Customer shall describe their approach to qualification and acceptance tests. This plan is intended to outline the Customer’s overall test philosophy along with an overview of the system-level environmental testing that will be performed to demonstrate the adequacy of the Auxiliary Passenger for ground and flight loads (e.g., static loads, vibration, acoustics, random and shock). The test plan shall include test objectives and success criteria, test specimen configuration, general test methods, and a schedule. It shall not include detailed test procedures. • An Auxiliary Passenger environment test file comprising theoretical analysis and test results following the system-level structural load and dynamic environment testing. This file should summarize the testing performed to verify the adequacy of the Auxiliary Passenger structure for flight and ground loads including venting. For structural systems not verified by test, a structural loads analysis report documenting the analyses performed and resulting margins of safety shall be provided. Arianespace may request to attend environmental tests for real time discussion of notching profiles and tests correlations. The final S/C status of compatibility shall be presented at the RAV launcher review.

Compatibility to launch system interfaces: As described in chapter 2.6.1., a mechanical and electrical fit check can be performed, prior to the initiation of the launch campaign. The conclusion of this fit-check will be presented at the RAV.

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6.3.4 Post-launch analysis After the flight, confirmation of the Auxiliary Passenger physical separation, obtained orbit and attitude at separation will be provided to the Customer at a time which is mission dependent (from one to several hours after separation depending on the available ground stations). Arianespace requires the Customer to provide satellite orbital tracking data on the initial Auxiliary Passenger orbit, including attitude just after separation.

6.4 Launch Vehicle adaptation Arianespace will supply the hardware and software to carry out the mission, complying with the Interface Control Document (DCI): • One adapter with separation system and, when necessary, umbilical interface connector(s); • One payload compartment using either the ASAP-S system on Soyuz or the VESPA system on Vega. • One mission logo installed on the fairing, based on the artwork supplied by Customer; • When necessary, one rack compatible with the launch pad installation, for the Check- Out Terminal Equipment (COTE).

6.5 Launch campaign 6.5.1 Typical Auxiliary Passenger launch campaign The Auxiliary Passenger launch campaign main phases are: • Preparation; • Arrival in French Guiana & transfer to the launch site; • Auxiliary Passenger autonomous preparation in PPF: Auxiliary Passenger preparation and checkout; • Auxiliary Passenger autonomous preparation in HPF (when necessary): Auxiliary Passenger hazardous operations; • Combined Operations (S/C mating on adaptor, Stack integration, fairing encapsulation, mating on launcher); • Launch countdown; • Transfer back of the GSE. All the facilities (PPF and HPF Clean Rooms and Lab for check-out stations - LBC) are shared with the other auxiliary payload(s) on the same flight.

6.5.1.1 Preparation phase & Operational documentation During the launch campaign preparation phase, to ensure activity coordination and compatibility with CSG facility, Arianespace issues the following operational documentation based on Application to Use Arianespace's Launch Vehicles and the Auxiliary Passenger Operations Plan (POS "Plan des Opérations Satellite"): • An Interleaved Operation Plan (POI); • A Combined Operations Plan (POC); • The set of detailed procedures for combined operations;

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• A countdown manual. The operational documentation and related items are discussed at technical meetings and status of the activity presented at final launch campaign preparation meeting. When deemed necessary, as an option available at customer request, Arianespace can organize a CSG visit for Auxiliary Passenger(s) Operations Plan preparation. It may comprise the visit of the facilities allocated to the auxiliary passenger(s), and review of operational documentation.

6.5.1.2 Spacecraft, GSE and propellant arrival & transfer to EPCU The Spacecraft, the ground support equipment and the propellant, if any, can be delivered to CSG by aircraft or by ship. The possible arrival areas are: • The Félix Eboué international airport for freighter airplanes and regular passenger flights. Small freight can be shipped by the regular Air France flights. The height of the container shall be lower than 1,58 m to be compatible with the fret doors of these passenger planes. • The Cayenne international harbor . • The Pariacabo docking area for Arianespace’s ships which ensure regular LV transport. These ships are also available for transferring spacecraft and/or GSE from Europe to CSG (option available at customer request). The usual stop harbors in Europe for containers loading are Rotterdam and Le Havre. Arianespace takes charge of the containers after their unloading from the plane or ship and ensures the transfer to the CSG by road.

6.5.1.3 Auxiliary Passenger autonomous preparation Autonomous operations and checks of the Auxiliary Passenger are carried out in the PPF. The Clean Rooms and Lab for check-out stations (LBC) is shared with the other auxiliary payload(s) on the same flight. The allocated areas in PPF and in LBC (as well as the number of allocated offices) will be provided according to the contractual provision. The activities shall be performed nominally within normal CSG working hours. Normal working hours at the CSG are based on 2 shifts of 8 hours per day, between 6:00 am and 10:00 pm from Monday to Saturday. After S/C departure from PPF (to HPF or UCIF), the evacuation of Ground Support Equipment from the clean room shall be completed within 1 working day.

6.5.1.4 Auxiliary Passenger hazardous operations (if necessary) When necessary, the auxiliary passenger is transferred in HPF for hazardous operations. After S/C departure from HPF, the evacuation of the GSE from the filling hall shall also be completed within 1 working day.

6.5.1.5 Combined Operations The Auxiliary Passenger shall be made available to Arianespace for the Combined Operations with the Launch Vehicle 13 (TBC) working days prior to the Launch, at the latest. The actual date will be defined in the Combined Operations Plan (POC) approved by the Customer. Similarly, the representative dummy shall be available for the combined operations readiness review (BT POC) at the latest.

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After delivery, verification and acceptance of all the upper composite parts, the BT POC authorizes the combined operations. The first combined operation is the Auxiliary Passenger integration on its adapter. It may occur in the PPF or in the HPF (when necessary). It is followed by the matting of the S/C with its adapter onto the carrying structure. The stack is then transferred to the Upper Composite Integration Facility (UCIF) for the matting of the main passenger and encapsulation. Depending on upper part configuration and of the launcher and, the UCIF can be S3B (for Soyuz) or S5A or S5B.

6.5.1.6 Launch pad operations The Auxiliary Passenger check-out equipment and specific COTE (Check Out Terminal Equipment) necessary to support the Auxiliary Passenger/Launch Vehicle on-pad operations shall be made available to ARIANESPACE, and validated, two days prior to operational use according to the approved operational documentation, at the latest. After the launch, all Auxiliary Passenger mechanical & electrical support equipment shall be removed from the various EPCU buildings & Launch Pad, packed and made ready for return shipment within three working days after the Launch.

6.5.2 Summary of launch campaign meetings and reviews

6.5.2.1 Auxiliary Passenger preshipment review Arianespace will be invited to the preshipment or equivalent review, organized by the customer and held before shipment of the Auxiliary Passenger to the CSG. The main outcomes of the preshipment review shall be communicated to Arianespace.

6.5.2.2 Satellite transport meeting Arianespace will hold a preparation meeting with the customer at the CSG just before satellite arrival. The readiness of the facilities at entrance port, and at CSG for satellite arrival, as well as status of formal issues and transportation needs will be verified.

6.5.2.3 Combined operations readiness review (BT POC “Bilan Technique POC") The objective of this review is to demonstrate the readiness of the Auxiliary Passenger, the flight items and the CSG facilities to start the combined operations according to POC. It addresses the following main points: • POC presentation, organization and responsibility for combined operations; • The readiness of the Upper composite items (adapters, carrying structure, fairing, and for a Soyuz launch, Fregat upper stage); • The readiness of the CSG facilities; • The readiness of the Auxiliary Passenger; • The availability of the Auxiliary Passenger representative dummy; • The mass of the payload in its final launch configuration.

6.5.2.4 Launch readiness review (RAL “Revue d’Aptitude au Lancement”) A Launch Readiness Review is held one or two days before launch and after the launch rehearsal. It authorizes the filling of the LV stages and the pursuit of the final countdown and launch. This review is conducted by Arianespace.

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6.5.2.5 Post flight debriefing (CRAL " Compte-rendu Après Lancement ") The day after the launch, Arianespace draws up a report to the Customer, on post flight analysis covering flight event sequences, evaluation of LV performance, and if available, injection orbit and accuracy parameters.

6.5.3 Range Support Arianespace provides a number of standard services and standard quantities of fluids described hereafter).

6.5.3.1 Transport services

Spacecraft, GSE and propellant transportation: Transport from and to one of the arrival areas and CSG at arrival and departure. Performed nominally within normal CSG working hours. And subject to advance notice. It does not include: • Unloading from plane or ship which is customer responsibility; • The “octroi de mer” tax on equipment permanently imported to Guiana, if any; • Insurance for spacecraft and its associated equipment.

Logistics support: Support for shipment and customs procedures for the spacecraft and its associated equipment and for personal luggage and equipment transported as accompanied luggage.

Spacecraft and GSE Inter-Site Transportation: All spacecraft transportation either inside the S/C container, and spacecraft GSE transportation between CSG facilities.

6.5.3.2 Payload preparation facilities allocation

PPF, HPF, LBC areas and meeting rooms: The Clean Rooms and Lab for check-out stations (LBC) is shared with the other auxiliary payload(s) on the same flight. The allocated areas in PPF (and HPF when necessary) and in LBC, as well as the number of allocated offices will be provided according to the contractual provision. The activities shall be performed nominally within normal CSG working hours. Storage: Any storage of equipment during the campaign. Two additional months for propellant storage, if any.

Schedule restrictions: The standard standalone launch campaign duration is limited to 10 working days for Micro Auxiliary Passenger and 15 working days for Mini auxiliary passenger. The Spacecraft shall be made available to Arianespace for the Combined Operations with the Launch Vehicle 13 (TBC) working days prior to the Launch, at the latest. The actual date will be defined in the Combined Operations Plan (POC) approved by the Customer.

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After S/C departure from PPF (to HPF or UCIF), the evacuation of Ground Support Equipment from the clean room shall be completed within 1 working day. Similarly, after S/C departure from HPF, the evacuation of the GSE from the filling hall shall also be completed within 1 working day. After the Launch, Spacecraft ground support equipment must be packed and removed from all EPCU facilities within maximum 3 working days.

6.5.3.3 Communication links The following communication services between the different spacecraft preparation facilities will be provided for the duration of a standard campaign (including technical assistance for connection, validation and permanent monitoring):

Service Type Quantity

RF- Link S/C/Ku/Ka band 1 TM / 1 TC through optical fiber Baseband Link S/C/Ku/Ka band 2 TM / 2 TC through optical fiber Data Link V11 and V24 network For COTE monitoring & remote control Ethernet Planet network, 10 Mbits/sec 1 VLAN available per project Umbilical Link Copper lines Up to 12 lines for Micro Auxiliary Passenger Up to 34 lines for Mini Auxiliary passenger Closed Circuit TV As necessary Intercom System As necessary Paging System 2 beepers per Project CSG Telephone As necessary Video Conference Equipment shared with other å Customers Internet Connection to local provider Note: å traffic to be paid, at cost, on CSG invoice after the campaign.

6.5.3.4 Cleanliness monitoring Continuous monitoring of organic deposit in clean room, with one report per week. Continuous counting of particles in clean room, with one report per week.

6.5.3.5 Fluid and gases deliveries

Gases Type Quantity

Compressed air Industrial, dedicated local As necessary network

GN2 N50, dedicated local network As necessary available at 190 bar

GN2 N30, dedicated network in S3 As necessary available at 190 area bar

GHe N55, dedicated local network As necessary, available at 410, 350 or 200 bar.

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Fluid Type Quantity LN2 N30 As necessary IPA MOS-SELECTIPUR As necessary Water Dematerialized As necessary

6.5.3.5 Safety equipment

Equipment Type Quantity Safety equipment for hazardous Standard As necessary operations (safety belts, gloves, shoes, gas masks, oxygen detection devices, propellant leak detectors, etc.)

6.6 Safety assurance 6.6.1 General The safety objectives are to protect the staff, facility and environment during launch preparation, countdown and flight. This is achieved through preventive and palliative actions: • Safety analysis based on the spacecraft safety submission; • Safety constraints during hazardous operations, and their monitoring and coordination; • Training and prevention of accidents; • Coordination of the first aide in case of accident; • For flight, short and long range flight safety analysis based on trajectory ground track. CSG is responsible for the implementation of the Safety Regulations and for ensuring that these regulations are observed. All launches from the CSG require approvals from Ground and Flight Safety Departments. These approvals cover payload hazardous systems design, all transportation and ground activities that involve spacecraft and GSE hazardous systems, and the flight plan.

6.6.2 Safety submission In order to obtain the safety approval, a Customer has to demonstrate that his equipment and its utilization comply with the provisions of the Payload Safety Handbook CSG-NT-SBU-16687-CNES. Safety demonstration is accomplished through several submission phases of documents defining and describing hazardous elements and their processing. The hazardous items check list is given in Annex 1 to help the customer for the establishment of the submission files. The typical time schedule for safety submissions is shown below:

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Safety submissions Typical schedule

Phase 0 – Feasibility (optional) Before contract A Customer willing to launch a spacecraft containing signature innovating systems can obtain a safety advice from CSG through the phase 0 submission.

Phase 1 - Design After the contract The submission of the spacecraft and GSE design and signature and before description of their hazardous systems. It shall cover PMA kick-off component choice, safety and warning devices, fault trees for catastrophic events, and in general all data enabling risk level to be evaluated.

End of Phase 1 submission No later than L - 12 m Phase 2 – Integration and qualification No later than L - 12 m The submission of the refined hardware definition and respective manufacturing, qualification and acceptance documentation for all the identified hazardous systems of the spacecraft and GSE. The submission shall include the policy for test and operating all systems classified as hazardous. Preliminary spacecraft operations procedures should also be provided.

End of Phase 2 submission No later than L - 7 m Phase 3 – Acceptance tests and hazardous operations L - 6 m The submission of the final description of operational procedures involving the spacecraft and GSE hazardous systems as well as the results of their acceptance tests if any.

Approval of the spacecraft compliance with CSG Before S/C fuelling at Safety Regulation and approbation of the procedures latest for autonomous and combined operations.

6.6.3 Safety measures during hazardous operations

The spacecraft authority is responsible for all spacecraft and associated ground equipment operations. The CSG safety department representatives monitor and coordinate these operations for all that concerns the safety of the staff and facilities. Any activity involving a potential source of danger is to be reported to the CSG safety department representative, which in return takes all measures necessary to provide and operate adequate collective protection, and to activate the emergency facilities. Each member of the spacecraft team must comply with the safety rules regarding personal protection equipment and personal activity. The CSG safety department representative permanently verifies their validity and he gives the relevant clearance for the any hazardous operations.

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On request from the Customer, the CSG can provide specific protection equipment for members of the spacecraft team. In case the launch vehicle, the spacecraft and, if applicable its co-passenger imposes crossed safety constraints and limitations, the Arianespace representatives will coordinate the respective combined operations and can restrict the operations or access to the spacecraft for safety reasons.

6.6.4 Safety training The general safety training will be provided through video presentations and documents submitted to the Customer before or at the beginning of the launch campaign. At the arrival of the launch team at CSG a specific training will be provided with on-site visits and detailed practical presentations that will be followed by personal certification. In addition, specific safety training on the hazardous operations, like fueling, will be given to the appointed operators, including operations rehearsals.

6.7 Quality assurance To achieve the highest level of reliability and schedule performance, the Arianespace’s Quality Assurance system covers the launch services provided to Customer, and extends up to the launch vehicle hardware development and production by major and second level suppliers. Arianespace quality rules and procedures are defined in the company’s Quality Manual. This process has been perfected through a long period of implementation, starting with the first Ariane launches more than 35 years ago, and is certified as compliant with the ISO 9000 standard. The system is based on the following principles and procedures: A. Appropriate management system The Arianespace organization presents a well-defined decisional and authorization tree including an independent Quality directorate responsible for establishing and maintaining the quality management tools and systems, and setting methods, training, and evaluation activities (audits). The Quality directorate representatives provide un-interrupted monitoring and control at each phase of the mission: hardware production, spacecraft-Launch vehicle compliance verification and launch operations. B. Configuration management, traceability and proper documentation system Arianespace analyses and registers the modifications or evolutions of the system and procedures, in order not to affect the hardware reliability and/or interfaces compatibility with spacecraft. The reference documentation and the rigorous management of the modifications are established under the supervision of the configuration control department. C. Quality monitoring of the industrial activities In complement to the supplier’s product assurance system, Arianespace manages the production under the following principles: acceptance of supplier’s Quality plans with respect to Arianespace Quality management specification; visibility and surveillance through key event inspection; approbation through hardware acceptance and non- conformance treatment.

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6.8 Optional Services The following Optional items and Services list is an abstract of the "Tailored and optional services list" available for the Customer:

System engineering: • Set of Preliminary mission analyses (as necessary depending of Auxiliary Passenger mission and characteristics) and Preliminary Mission Analysis Review.

Interface tests: • Fit-check (mechanical/electrical) with flight adapter in CSG; • Fit-check (mechanical/electrical) with representative flight adapter at Customer's premises; • Fit-check (mechanical/electrical) with representative flight adapter hardware at Customer's premises, followed by shock/drop test.

Compatibility tests: • Shock test device to perform pyro shock test (recommended when using the SSASAP5 adapter to demonstrate compatibility to shock environment generated by the SSASAP5 adapter at S/C separation); • VESTA test (recommended for a Vega launch to demonstrate compatibility to shock environment generated by LV events).

Range Operations: • CSG site survey; • Spacecraft and/or GSE transport to Kourou, using an Arianespace ship to transport the spacecraft and/or its associated equipment and propellant; • Extra working shift; • Campaign extension above contractual duration; • Chemical analysis (gas, fluids and propellants except Xenon); • S/C weighting.

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APPLICATION TO USE ARIANESPACE’S LAUNCH VEHICLE (DUA) Annex 1 FOR AUXILIARY PASSENGERS

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The Customer interested in a launch opportunity as auxiliary passenger for a small spacecraft shall provide to ARIANESPACE the information described in the present annex.

The following Application to Use Arianespace’s Launch Vehicle (DUA) template, will preferably be provided, duly completed, in MS Word format, along with a Gantt-chart of S/C preparation schedule, a CAD model (*.stp format) and all relevant electronic files (MS Excel). A more detailed updated version of the DUA might be provided after signature of the LSA, along with FEM and thermal models.

A1.1. Spacecraft description and mission summary

Manufactured by : TBD Platform type : TBD

DESTINATION Earth Observation* Scientific* Meteorological* Navigation* Telecommunication* In Orbit Test/Demonstration* Others*

MASS LIFETIME Total mass at launch TBD kg TBD years

OPERATIONAL ORBIT DIMENSIONS a × e × inclination; ω; RAAN Stowed for launch H TBD mm L TBD mm W TBD mm PREFERRED INJECTION ORBIT Deployed on orbit a × e × inclination; ω; RAAN H TBD mm L TBD mm W TBD mm PAYLOAD Purpose & brief description of the instrument(s)

COMMUNICATION SUB-SYSTEM Frequency band for TM &TC, number of receivers/antennas and location PROPULSION SUB-SYSTEM Brief description: chemical/electrical prop. system, type of propellant, number of tanks, number of thrusters,...

ELECTRICAL POWER SUB-SYSTEM Solar array description (L x W) Beginning of life power TBD W End of life power TBD W Batteries description TBD (type, capacity)

ATTITUDE CONTROL SUB-SYSTEM Brief description: sensors description (Sun, Stellar, …), actuators description ( momentum wheels, thrusters, …)

GROUND STATION NETWORK For LEOP phase: TBD For operational phase: TBD

Note : * to be selected.

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A1.2.1 Launch period Provide targeted launch period/launch slot.

A1.2.2 S/C main milestones Provide a Gantt chart of the S/C design, manufacturing and tests schedule with the following main milestones: - System PDR, - System CDR, - Start/end of manufacturing for each main S/C parts (platform, instruments, …), - Start/end of each main part integration, - Start/end of S/C integration, - Start/end of S/C test campaign, - Flight acceptance review (FAR).

A1.2.3 Contents of the spacecraft development plan The Customer will prepare a file containing all the documents necessary to assess the spacecraft development plan with regard to the compatibility with the launch vehicle. It shall include, at least: - spacecraft test plan: define the qualification policy, vibrations, acoustics, shocks, protoflight or qualification model, - tests configuration (S/C representativeness, tests adapter, etc…), - test facility location (Customer’s or Manufacturer’s facility), - if any, necessary additional tests at the range.

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A1.3. Mission characteristics A1.3.1 Orbit description Indicate preferred injection orbit parameters and, if different, the Spacecraft operational orbit. Indicate the acceptable orbit dispersions (at 3 σ).

Injection orbit at S/C S/C operational orbit separation (when relevant) Semi major axis a _____ ±_____ km ______km Eccentricity e _____ ±______Inclination i _____ ±_____ deg ______deg Argument of perigee ω _____ ±_____ deg ______deg Right Ascension of _____ ±_____ deg ______deg Ascending Node RAAN

A.1.3.2 Launch time / window For SSO mission, provide the preferred Local Time of Ascending Node (LTAN). For any other orbit, provide the preferred launch window (preferably in an electronic file, MS Excel). Constraints on opening and closing shall be identified and justified.

A1.3.3 Flight and separation conditions

A1.3.3.1 Separation conditions Separation mode and conditions Indicate preferred separation mode (3-axis stabilized, low axial or transverse spin, etc...). Indicate acceptable depointing, tip-off rates and relative velocity at separation. Separation attitude Indicate the preferred orientation at separation. For circular or nearly circular orbits, the desired orientation at separation should be specified by the Customer with respect to the following inertial reference frame [U, V, W] related to the orbit at S/C separation time, as defined below: U = Radius vector with its origin at the center of the Earth, and passing through the intended separation point. V = Vector perpendicular to U in the intended orbit plane, having the same direction as the orbit velocity. W = Vector perpendicular to U and V to form a direct trihedron (right-handed system [U, V, W]). For 3-axis stabilized separation mode, two of the three S/C axes [U, V, W] coordinates should be specified.

A1 - p68 Arianespace ©, June 2017 DUA for Auxiliary Passengers Auxiliary Passengers User’s Manual, Issue 1 A1.3.3.2 Attitude during ascent phase, prior to S/C separation If any, indicate any particular S/C attitude limitation (solar aspect angle constraints, spin limitation, etc…), applicable during the ascent phase and/or during the coast phases.

A1.3.3.3 Other conditions If any, indicate any other S/C limitations including: - maximum acceptable aerothermal flux, - any flight duration limitation, - any constraints for ground station visibility, - etc…

A1.3.4 Sequence of events after S/C separation Describe the sequence of events after the S/C separation from the launcher, including: - on-board Computer switch-on, - TM emitters switch-on, - propellant system priming, - attitude Control System switch-on, - any deployments (solar generators, booms, etc…), - etc…

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A1.4. Spacecraft description

A1.4.1 Spacecraft systems of axes Provide a description of spacecraft system of axes (please, include a sketch). The origin of the axes shall be in the mounting plane. The axes are noted Xs, Ys, Zs and shall form a right handed trihedron. All the S/C data and models shall be given considering the same spacecraft system of axes, including S/C mass properties, CAD model, FEM model, etc…

A.1.4.2 Spacecraft geometry in the flight configuration Provide a CAD model (*.stp format) of the spacecraft in flight configuration together with the associated drawings. Additionally, provide: - detailed drawings of the interface with adapter, with manufacturing tolerances; - detailed dimensional data (including manufacturing tolerances, any MLI, electrical harness, …) for the S/C critical elements, that is the S/C closest parts to the fairing, carrying structure and adaptor: solar array panels, deployment mechanisms, etc....

A1.4.3 Spacecraft mass properties Provide the S/C nominal mass properties and associated dispersion (Min/Max) in launch configuration.

C of G Mass Coefficients of inertia Matrix coordinates (kg) (kg. m2) (mm)

M XG YG ZG Ixx Iyy Izz Pxy Pyz Pzx Nominal Tolerance Min/Max Min/Max Min/Max Min/Max Min/Max Min/Max

Notes: - Center of Gravity coordinates are referenced in the spacecraft coordinate system. The origin is the geometrical center of the separation plane. - Moments of Inertia are referenced in the spacecraft coordinate system where the origin is at the Center of Gravity of the spacecraft. - Products of Inertia are calculated by the following equation: Pxy = + ∫xy dm.

In the case the adaptor is supplied by the Customer, provide also mass properties of spacecraft with adapter, and mass properties of adapter alone just after separation.

A1 - p70 Arianespace ©, June 2017 DUA for Auxiliary Passengers Auxiliary Passengers User’s Manual, Issue 1 A1.4.4 Fundamental modes Indicate fundamental modes (lateral, longitudinal) of spacecraft hardmounted at interface.

A1.4.5 Propellant/pressurant characteristics Provide the propellant and pressurant tanks description, and if relevant, propellant sloshing characteristics:

Propellant tanks #1 … Propellant Density (kg/m 3) Tank volume (l) Fill factor (%) Liquid volume (l) Liquid mass (kg) Center of gravity Xs of propellant Ys loaded tank Zs Pendulum mass (kg) Pendulum length (m) Pendulum Xs attachment Ys point Zs Fixed mass (if any) Slosh model Fixed mass Xs under 0 g attachment Ys

point (if any) Zs

Natural frequency of fundamental sloshing mode (Hz) Pendulum mass (kg) Pendulum length (m) Pendulum Xs attachment Ys point Zs Slosh model Fixed mass (if any) under 1 g Fixed mass Xs

attachment Ys

point (if any) Zs Natural frequency of fundamental sloshing mode (Hz)

Pressurant Tanks #1 … Pressurant Volume (l) Loaded mass (kg) Xs Center of gravity (mm) Ys Zs

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Arianespace proposes a series of standard adapters for Small Satellites. Depending of the size, the mechanical interface can be a bolted I/F, in that case both the so-called passive ring and active ring is provided by Arianespace, or a clamp band I/F, in that case the active ring only is provided by Arianespace. Interface geometry:

Provide a drawing with detailed dimensions and nominal tolerances showing:

For Small-sat with clamp band I/F: - The spacecraft interface ring; - The area allocated for spring actuators and pushers; - The area allocated for microswitches; - Umbilical connector locations and supports; - Any equipment in close proximity to the separation clamp band (thrusters, antennas, MLI, etc…). For Small-sat with bolted I/F: - The spacecraft rear panel; - Umbilical connector locations and supports; - Any equipment in close proximity to the separation plane (thrusters, antennas, MLI, etc…).

Interface material description: For each spacecraft mating surface in contact with the launcher adapter (and the clampband, when relevant) indicate material, roughness, flatness, surface coating, rigidity (frame only), inertia and surface (frame only) and grounding.

A1.4.7 Electrical interfaces

Provide the following: - The location of the spacecraft ground potential reference on the spacecraft interface frame; - Data link requirements on ground (baseband and data network) between spacecraft and EGSE; - Definition of umbilical connectors and links in a table form (preferably in an electronic file, MS Excel):

S/C connector Function Max voltage (V) Max current Expected one pin allocation (mA) way resistance number (Ω) 1 2 3 …

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A1.4.8.1 S/C Telecommunication sub-system(s) general description Provide the S/C Telecommunication system(s) main characteristics: - description of S/C telemetry (TM) and telecommand (TC) systems; - description of TM et TC antennas, antenna location, and antenna pattern; - for information, brief description of payload telecommunication system(s).

A1.4.8.2. Spacecraft ground station network Provide the list of ground station to be used for spacecraft acquisition and early operations after S/C separation from the launcher.

A1.4.8.3 Spacecraft telemetry (TM) and telecommand (TC) systems Provide a detailed description of spacecraft telemetry (TM) and telecommand (TC) systems (preferably in an electronic file, MS Excel):

Source unit designation Tx1 Tx… Rx1 Rx… Function Band

Carrier Frequency, F 0 (MHz) Bandwidth centered -3 dB

around F 0 -20 dB -60 dB Carrier Modulation Type Index Bit rate Sub Carrier (MHz) Minimum S/N (dB) associated bandwidth (MHz) Local Oscillator Frequency (MHz) 1st intermediate Frequency (MHz) 2nd intermediate Frequency (MHz) Field strength at antenna, receive Max 2 (dBW/m ) Nom Min RF Output Impedance (Ohm) Lower Power mode availability (Yes/no) Antenna designation Horn Omni Horn Omni Antenna Type Location X,Y,Z Pattern Gain max (dBi) EIRP: Output power (dBW) Max Nom Min Antenna Input power (dBW) Max Nom Min

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If any, provide the radio link needs between spacecraft, spacecraft check-out system and PPF facility. Provide the spacecraft transmission plan as shown in table below:

Source unit description Tx1 Tx... Rx1 Rx… Function TBD TBD During preparation on launch site (PPF) TBD TBD During HPF activities OFF OFF Countdown before H0-1H30mn OFF OFF After H0–1H30mn until TBD s after OFF OFF separation* In orbit (or in transfer orbit) TBD TBD * Actual delay will be determined in the frame of mission analysis.

A1.4.9. Other S/C characteristics

Provide any other S/C characteristics and/or limitations, if any, including:

- If any, contamination constraints and contamination sensible surfaces; - Maximum ascent depressurization rate and differential pressure; - Temperature and humidity limits during launch preparation and flight phase; - If available, S/C electrical field susceptibility levels and S/C sensitivity to magnetic fields.

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A1.5.1 Provisional range operations schedule Provide list of main operations, with description and estimated timing. Identify all hazardous operations.

A1.5.2 Facility requirements For each facility used for spacecraft preparation (PPF, HPF) provide: - Main operations list and description - Surface area needed for spacecraft, GSE and Customer offices - Environmental requirements (Temperature, relative humidity, cleanliness) - Power requirements (Voltage, Amps, # phases, frequency, category) - RF and hardline requirements - Support equipment requirements - GSE and hazardous items storage requirements

A1.5.3 Communication needs For each facility used for spacecraft preparation (PPF, HPF) provide need in telephone, facsimile, data lines, time code etc.

A1.5.4 Handling, dispatching and transportation needs Provide: - Estimated packing list with indication of designation, number, size (L x W x H in m) and mass (kg) - Propellant transportation plan (including associated paperworks) - A definition of the spacecraft container and associated handling device (constraints) - A definition of the spacecraft lifting device - In case the adapter is provided by the customer, a definition of adapter interface - A definition of spacecraft GSE (dimensions and interfaces required) - Dispatching list

A1.5.5 Others

A1.5.5.1 List of fluids Indicate type, quality, quantity and location for use of fluids to be supplied by Arianespace.

A1.5.5.2. Chemical and physical analysis to be performed on the range Indicate for each analysis: type and specification.

A1.5.5.3. Safety garments needed for propellants loading Indicate number.

A1.5.5.4. Technical support requirements Indicate need for workshop, instrument calibration.

A1.5.5.5. Security requirements If any, provide specific security requirements.

Arianespace ©, June 2017 A1 - p75 DUA for Auxiliary Passengers Auxiliary passengers User’s Manual Issue 1 A1.5.6. Documentation: Contents of Spacecraft Operations Plan (POS) The Customer will be asked to provide a Spacecraft Operations Plan which will define the operations to be executed on the spacecraft from arrival at the CSG, at the launch site, and up to the launch. A typical content is presented here below: 1. General 1.1 Introduction 1.2 Applicable documents 2. Management 2.1 Time schedule with technical constraints 3. Personnel 3.1 Organizational chart for spacecraft operation team in campaign 3.2 Spacecraft organizational chart for countdown 4. Operations 4.1 Handling and transport requirements for spacecraft and ancillary equipment 4.2 Tasks for launch operations (including description of required access after integration on carrying structure and/or fairing encapsulation) 5. Equipment associated with the spacecraft 5.1 Brief description of equipment for launch operations 5.2 Description of hazardous equipment (with diagrams) 5.3 Description of ground equipment (when in PPF, HPF, and Launch Pad) 6. Installations 6.1 Surface areas 6.2 Environmental requirements 6.3 Communications 7. Logistics 7.1 Transport facilities 7.2 Packing list

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A1.6 Safety aspects

A1.6.1. S/C hazardous systems and operations Provide a list of: - the S/C hazardous system (propellant, electro-pyrotechnic devices, batteries, laser, ionizing sources, etc…) - the intended hazardous activities for S/C preparation during S/C launch campaign at CSG (S/C handling, propellant loading, battery charging, deployment tests, etc…)

A1.6.2. Safety submission The Customer will be asked to provide Safety files for safety submissions, according to Payload Safety Handbook CSG-NT-SBU-16687-CNES. These files will contain a description of the hazardous systems and operations and will respond to all questions on the hazardous items check list given in the Payload Safety Handbook here below:

A1 Solid-propellant engine A2 Ignition module, safe and arm unit, command and control circuits A3 Corresponding ground segment equipment and operations B1 Electro-pyrotechnic devices - Compliance B2 Command and control circuit B3 Corresponding ground segment equipment and operations C1 Monopropellant propulsion system C2 Valve command and control circuit C3 Corresponding ground segment equipment and fuelling equipment AC1 Bipropellant propulsion system AC2 Valve command and control circuit AC3 Corresponding ground segment equipment and fuelling equipment D1A Non-ionizing radiation D2A Optical systems D3A Lasers D1B Batteries and electrical systems D2B Command and control D3B Corresponding ground segment equipment D1C Fluids and gases other than propellant – Cryogenic products D2C Command and control D3C Corresponding ground segment equipment D1D Mechanical and electromechanical equipment, structures, transport and handling equipment D2D Equipment and other systems D1E Ionizing radiation – Flight sources D3E Ionizing radiation – ground segment equipment O Documentation GC Miscellaneous

A1.7 Miscellaneous Provide any other specific requirements for the mission or S/C preparation. Provide a list of acronyms and symbols with their definition.

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STANDARD AUXILIARY PASSENGERS ADAPTERS Annex 2

1.1. . Several standard off-the-shelf adapters are available ensuring interfaces between the launcher and the micro or mini spacecraft.

Beside the historical SSASAP5 ring from Dassault which provides a bolted interface at Ø 298 mmm, a family of low-shock adapters from RUAG Space Company are available for Micro and Mini Auxiliary Passengers, as well as a light low-shock 937 adapter from Airbus DS for Mini Auxiliary Passengers.

The SSASAP5 ring from Dassault has been widely and successfully used on many Ariane launches and more recently on Vega and Soyuz. More than 45 units have flown so far since 1988. The separation system is composed of a pyro-cutting system and up to 10 pushers.

The new family of low shock adapters from RUAG Space Company uses a down-scaled version of the existing clamp band and CBOD of the larger adapters used for Ariane and Soyuz passengers (PAS 937 S, PAS 1194 VS, etc…). They consist of the adapter structure, the clamp band assembly together with its bracket set, the separation spring set and umbilical bracket attached to the structure. The available clamp band diameters are 381, 432 and 610 mm and the LV – S/C interface can be either bolted (as for the SSASAP5 ring) or at the clampband (as for larger S/C).

For the 937 mm diameter, the Active Ring 937 (AR 937) from Airbus DS, is available. It consists of an aluminum alloy conical structure, clamp band (LPSS 937) assembly, bracket set, separation spring set and a protective membrane.

The table below summarizes the available options.

Interface at the clamp Bolted interface band SSASAP5 X (Ø298 mm) PAS 381 S X ( Ø381 mm) PAS 432 S X (Ø298 mm) X (Ø432 mm) PAS 610 S X (Ø610 mm) AR 937 X (Ø945 mm)

In case of bolted interface, a part of the adapter (the so-called passive ring) remains on the S/C after separation.

The adapters hold the electrical harness that is necessary for umbilical links as well as for separation orders and telemetry data transmission. This harness will be tailored to user needs, with its design depending on the required links between the spacecraft and the launch vehicle (see Chapter 2).

In some situations, the Customer may wish to assume responsibility for payload adapter. In such cases, the Customer shall ask the Arianespace approval and corresponding requirements. Arianespace will supervise the design and production of such equipment to insure the compatibility at system level.

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A2.1 SSASAP5

The SSASAP5 ring is suitable for micro Auxiliary Passenger up to 120 kg.

The micro S/C is mounted on the SSASAP5 ring via 12 Ø6 bolts (and the SSASAP5 ring is mounted on the LV carrying structure - ASAP-S or VESPA - via 12 Ø8 bolts) provided by Arianespace. The required torque of the SSASAP5 attachment bolts into the base of the spacecraft is 0.8 mdaN.

The separation is triggered by two pyro detonators which are initiated by a redundant command and which allow to cut the SSASAP5 structure. The corresponding shock environment (Flight Limit Loads) is presented in Figure A2.1.2.

The SSASAP5 separation system provides a relative velocity along the spacecraft longitudinal axis, which is adjustable between 1 m/s and 3 m/s, the final value being defined by Arianespace, in accordance with the Auxiliary Payload customer. The velocity is adjusted by choosing the number of springs (up to 10) and the energy of the springs (adjustable by mean of props).

After separation, a residual mass of 1 kg remains on the micro spacecraft.

Figure A2.1.1: SSASAP5 - Separation system logic

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Figure A2.1.2: SSASAP5 - Shock spectrum at separation

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Figure A2.1.3: SSASAP5 mechanical interface for Micro Auxiliary Passenger

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Dimensions and tolerances A = 348 ± 0.1 mm (diameter) D = 6.5 ± 0.1 mm (diameter) B = F = 298 ± 0.1 mm (diameter) E = 8.5 ± 0.1 mm (diameter) C =264.3 ± 0.1 mm (diameter) G = 28 ± 0.1 mm

Figure A2.1.4: dimensions, tolerances and positioning of the separation system

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Figure A2.1.5: prohibited volume of micro auxiliary passenger

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Figure A2.1.6: Micro satellite mounting bolt definition

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A2.2 PAS 381 S

The PAS 381 S is designed and qualified to support a payload of 200 kg centered at 0.5 m from the separation plane.

The PAS 381 S is composed of two parts: - Spacecraft Ring Assembly (so-called passive ring) - PAF 381S (so-called active ring)

The Spacecraft Ring upper interface towards the spacecraft has a 381 mm diameter bolted interface with 24 holes for ¼-inch bolts.

The PAF 381 S itself is mainly composed of: • A monolithic aluminum structure with a diameter of 381 mm at the level of the separation plane • A clamp band assembly with a Low Tension Clamp Band Opening Device (CBOD-LT) • A set of actuators (4 to 24)

Clamp Band release is obtained thanks to a pyrotechnically initiated Low Tension Clamp Band Opening Device (CBOD-LT). The CBOD-LT is specially designed to generate low shock levels. The corresponding shock environment (Flight Limit Loads) is presented in Figure A2.2.1.

The clamp band pretension is 11 kN. A set of 4 catchers secures a safe behavior and parks the clamp band on the adapter.

The spacecraft is forced away from the launch vehicle by up to 24 actuators, bearing on supports fixed to the passive ring.

The typical mass of the PAF 381 S adapter system is 2.8 kg. The typical mass of the passive ring (remaining attached to the spacecraft) is 1.0 kg

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Figure A2.2.1: PAS 381 S - Shock spectrum at separation

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Figure A2.2.2: PAS 381 S – General view

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Figure A2.2.3: PAS 381 S – Clamp band assembly interface

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A2.3 PAS 432 S

The PAS 432 S is designed and qualified to support a payload of more than 400 kg centered at 0.5 m from the separation plane.

The PAS 432 S is composed of two parts: - Spacecraft Ring Assembly (so-called passive ring) - PAF 432S (so-called active ring) The Customer can use either the complete PAS 432S assembly, or the PAF 432S alone. In the latter case, the spacecraft rear frame shall precisely respect the mechanical I/F drawings detailed in the following page.

When the complete assembly is used, the Spacecraft Ring upper interface towards the spacecraft has a 298 mm diameter bolted interface with 12 holes for M8 bolts.

The PAF 432 S itself is mainly composed of: • A monolithic aluminum structure with a diameter of 432 mm at the level of the separation plane • A clamp band assembly with a Low Tension Clamp Band Opening Device (CBOD-LT) • A set of up to 6 actuators

Clamp Band release is obtained thanks to a pyrotechnically initiated Low Tension Clamp Band Opening Device (CBOD-LT). The CBOD-LT is specially designed to generate low shock levels. The corresponding shock environment (Flight Limit Loads) is presented in Figure A2.3.1.

The clamp band pretension is 15 kN and the corresponding maximum tension (during installation) is 18 kN. A set of 4 catchers secures a safe behavior and parks the clamp band on the adapter.

The spacecraft is forced away from the launch vehicle by up to 6 actuators, bearing on supports fixed to the passive ring (or to the spacecraft rear frame).

The force exerted on the spacecraft by each spring does not exceed 200 N.

The typical mass of the PAF 432 S adapter system is 5.3 kg. The typical mass of the passive ring (remaining attached to the spacecraft) is 1.9 kg

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Figure A2.3.1: PAS 432 S - Shock spectrum at separation

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Figure A2.3.2: PAS 432 S (Passive & active rings) – General view

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Figure A2.3.3: PAS 432 S (Passive & active rings) – Actuators

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Figure A2.3.4: PAS 432 S (Passive & active rings) – Umbilical connectors

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Figure A2.3.5: PAS 432 S (Passive & active rings) – Clamp band assembly interface

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Figure A2.3.6: PAF 432 S (Active ring only) – General view

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Figure A2.3.7: PAF 432 S (Active ring only) – Interface frames

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Figure A2.3.8: PAF 432 S (Active ring only) – Actuators

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Figure A2.3.9: PAF 432 S (Active ring only) – Umbilical connectors

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Figure A2.3.10: PAF 432 S (Active ring only) – Clamp band assembly interface

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A2.4 PAS 610 S

The PAS 610 S is designed and qualified to support a payload of 400 kg centered at 0.85 m from the separation plane.

The PAS 610 S is mainly composed of:

• A monolithic aluminum structure with a diameter of 610 mm at the level of the spacecraft separation plane • A clamp band assembly with a Low Tension Clamp Band Opening Device (CBOD-LT) • A set of up to 12 actuators

The spacecraft is secured to the adapter interface frame by the clamp band assembly. The clamp band consists of a band with one connecting point. The tension applied to the band provides pressure on the clamp which attaches the satellite to the launcher. Release is obtained thanks to a pyrotechnically initiated Low Tension Clamp Band Opening Device (CBOD-LT). The CBOD-LT is specially designed to generate low shock levels. The corresponding shock environment (Flight Limit Loads) is presented in Figure A2.4.1.

The clamp band pretension is 15 kN and the corresponding maximum tension (in flight) is 19.4 kN. A set of 5 catchers secures a safe behavior and parks the clamp band on the adapter.

The spacecraft is forced away from the launch vehicle by up to 12 actuators, bearing on supports fixed to the spacecraft rear frame.

The force exerted on the spacecraft by each spring does not exceed 230 N. If necessary, the stroke of each spring can be limited in order to tune the energy provided by each spring, allowing counteracting the effect of spacecraft nominal static unbalance at spacecraft separation.

The typical mass of the PAS 610 S adapter system is 6 kg.

After separation, there is no PAS 610 S part remaining on the spacecraft.

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Figure A2.4.1: PAS 610 S - Shock spectrum at separation

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Figure A2.4.2: PAS 610 S – General view

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Figure A2.4.3: PAS 610 S – Interface frames

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Figure A2.4.4: PAS 610 S – Actuators

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Figure A2.4.5: PAS 610 S – Umbilical connectors

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Figure A2.4.6: PAS 610 S – Clamp band assembly interface

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A2.5 AR 937

The Active Ring 937 (AR 937) is designed and qualified to support a payload up to 710 kg centered at 1.2 m from the separation plane.

The AR 937 is mainly composed of:

• A conical structure made of aluminum alloy with a diameter of 945 mm at the level of the spacecraft separation plane; • A Low-shock Payload Separation System (LPSS 937*light) with : o A Clamp band assembly that joins both the S/C and the adapter by means of a ring around the common interface of the two structures. o A Release device that triggers the opening of the band by pyrotechnics means and guides, parks, and catches the band ring to leave the S/C free to be ejected from the adaptor. • A Jettisoning device, consisting in a set of 4 to 8 springs, that provides the necessary energy to separate the two structures; • A protective membrane (for a Vega launch).

The LPSS 937*light design is an evolution of the LPSS* that has already flown several times on Ariane 5 and Vega. The Release device is designed to generate low shock levels. The corresponding shock environment (Flight Limit Loads) is presented in Figure A2.5.1.

The clamp band pretension is 24 kN and the corresponding maximum tension (in flight) is 30 kN. The release kinematics of the band is controlled by the 6 catcher slots (also called ¨guides¨).

The set of springs (4 to 8) is used to provide the necessary energy to impulse the S/C away from the carrying structure. Each spring has a maximum force of 450N and a maximum energy of 10J. Half energy can be achieved by the installation of dedicated special bushing and letting the 40mm stroke reduced.

The typical mass of the AR 937 adapter system is 45 kg.

After separation, there is no AR 937 part remaining on the spacecraft.

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Figure A2.5.1: AR 937 - Shock spectrum at separation

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Figure A2.5.2: AR 937 – General view

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Figure A2.5.3: AR 937 – Interface frames

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Figure A2.5.4: AR 937 – Actuators

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Figure A2.5.5: AR 937 – Umbilical connectors

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Figure A2.5.6: AR 937 – Clamp band assembly interface

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AUXILIARY PASSENGERS Annex 3 LAUNCH RECORD

LAUNCH LAUNCHER SATELLITE FLIGHT CUSTOMER COUNTRY MISSION MASS ORBIT DATE VERSION AMSAT P3C (Oscar 13) V-22 15/06/1988 AR 44LP AMSAT DE Germany TELECOM 139 GTO UoSAT-D AR 40 SSTL UK TELECOM + TECHNO 44,5 SSO

UoSAT-E AR 40 SSTL UK TELECOM + TECHNO 45,5 SSO

Pacsat (Oscar 14) AR 40 AMSAT NA USA TELECOM 12 SSO V-35 22/01/1990 DOVE (Oscar 15) AR 40 AMSAT BRAZIL Brazil TELECOM 11,6 SSO Webersat (Oscar 16) AR 40 AMSAT NA USA TELECOM + TECHNO 14,8 SSO

Lusat (Oscar 17) AR 40 AMSAT Argentina Argentina TELECOM 12,5 SSO

ORBCOMM-X AR 40 Orbcomm USA TELECOM 16 SSO SARA AR 40 ESIEESPACE France TELECOM + SCIENCE 17,5 SSO V-44 17/07/1991 TubSat A AR 40 Technische Universität Berlin Germany TECHNO 32,5 SSO

UoSAT-F AR 40 SSTL UK TELECOM + TECHNO 46,5 SSO

Kitsat A AR 42P KAIST South Korea TECHNO 46,1 LEO V-52 10/08/1992 S80/T AR 42P CNES France TELECOM 47,5 LEO

ARSENE V-56 12/05/1993 AR 42L RACE (Radio Amateur Club de l'Espace / CNES) France TELECOM 154 GTO

STELLA AR 40 CNES France SCIENCE 48 SSO Healthsat 1 AR 40 SatelLife USA TELECOM 42,3 SSO

Kitsat B AR 40 KAIST South Korea TECHNO 47,2 SSO V-59 26/09/1993 Eyesat A AR 40 Interferometrics Inc. / AMSAT NA USA TELECOM 10,5 SSO

Itamsat AR 40 AMSAT Italy Italy TELECOM 9,9 SSO PoSat 1 AR 40 POSAT Consortium Portugal TECHNO 47,8 SSO

STRV 1A AR 44LP UK Ministry Of Defence UK TECHNO 50,8 GTO V-64 17/06/1994 STRV 1B AR 44LP UK Ministry Of Defence UK TECHNO 52,3 GTO CERISE AR 40 DGA France ELINT 46,3 SSO V-75 07/07/1995 UPM/Sat 1 AR 40 Universidad Politécnica de Madrid Spain TECHNO 45,3 SSO

Equator-S V-103 02/12/1997 AR 44P Deutsche Agentur für Raumfahrtangelegenheiten Germany SCIENCE 229 GTO

Clémentine V-124 03/12/1999 AR 40 DGA France ELINT 46,1 SSO AMSAT P3D (Oscar40) AR 5 AMSAT DE Germany TELECOM 629 GTO

STRV 1C V-135 16/11/2000 AR 5 UK Ministry Of Defence UK TECHNO 105 GTO

STRV 1D AR 5 UK Ministry Of Defence UK TECHNO 103 GTO LDREX V-138 19/12/2000 AR 5 NASDA Japan TECHNO 183 GTO

Idefix V-151 04/05/2002 AR 42P AMSAT France France TELECOM 12,3 SSO

Smart 1 V-162 27/09/2003 AR 5G ESA Europe LUNAR EXPL. 367 GTO

Essaim 1 AR 5G DGA France ELINT 118 SSO Essaim 2 AR 5G DGA France ELINT 118 SSO

Essaim 3 AR 5G DGA France ELINT 118 SSO V-165 18/12/2004 Essaim 4 AR 5G DGA France ELINT 118 SSO PARASOL AR 5G CNES France EARTH OBS. 108 SSO

NANOSAT AR 5G INTA (Instituto Nacional de Técnica Aeroespacial) Spain TECHNO 19 SSO

LDREX 2 V-173 13/10/2006 AR 5ECA NASDA Japan TECHNO 211,9 GTO

SPIRALE A AR 5ECA DGA France EARLY WARNING 117,5 GTO V-187 12/02/2009 SPIRALE B AR 5ECA DGA France EARLY WARNING 117,5 GTO

SSOT SYZ ST Chilean Ministry of Defence Chile EARTH OBS. 117 SSO

ELISA 1 SYZ ST DGA France ELINT 120 SSO ELISA 2 VS 02 17/12/2011 SYZ ST DGA France ELINT 120 SSO

ELISA 3 SYZ ST DGA France ELINT 120 SSO

ELISA 4 SYZ ST DGA France ELINT 120 SSO

ALMASat 1 VEGA University of Bologna Italy TECHNO 12,5 SSO e-st@r VEGA Politecnico di Torino Italy TECHNO 1 LEO

Goliat VEGA University of Bucharest Romania TECHNO 1 LEO

MaSat 1 VEGA Budapest University of Technology and Economics Hungary TECHNO 1 LEO VV 01 13/02/2012 PW-Sat 1 VEGA Warsaw University of Technology Poland TECHNO 1 LEO

ROBUSTA VEGA University of Montpellier II France TECHNO 1 LEO

UniCubeSat-GG VEGA GAUSS (La Sapienza University of Rome) Italy TECHNO 1 LEO

XaTcobeo VEGA INTA, Universidade de Vigo Spain TECHNO 1 LEO VNREDSat-1 VEGA Vietnamese Academy of Science and Technology (VAST) Vietnam EARTH OBS. 115 SSO VV02 07/05/2013 ESTCube 1 VEGA National University of Tartu / Estonian Space Office Estonia TECHNO 1 SSO

MICROSCOPE SYZ ST CNES France SCIENCE 301,5 SSO OUFTI-1 SYZ ST Fly Your Satellite! Program: University of Liège Belgium TECHNO 1 SSO VS14 25/04/2016 E-st@r-II SYZ ST Fly Your Satellite! Program: Politecnico di Torino Italia TECHNO 1 SSO AAUSAT 4 SYZ ST Fly Your Satellite! Program: Aalborg University Denmark TECHNO 1 SSO

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