Rockot User's Guide

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Rockot User's Guide

EHB0003, Issue 5, Revision 0 August 2011

EUROCKOT Launch Services GmbH www.eurockot.com

Dear Customer, It is my pleasure to present to you issue 5 of the EUROCKOT User's Guide in succession to issue 4 which was published in 2004. A notable milestone occurred when our company performed its tenth Customer mission with the successful launch of SERVIS-2 in June 2010. The year 2010 also marked the tenth anniversary of becoming operational, recalling the first flight of the Rockot with the Breeze-KM upper stage in May 2000. I am also happy to report that our manifest will see launches continuing for the European Space Agency starting in 2012 with the joint launch of the three SWARM satellites, under- pinning our success as a leading provider of launch services in the small launcher market segment. We look forward to serving you, our Customer, with the reliability, flexibility and commercial benefits you have come to know during the course of this decade. EUROCKOT is the joint venture of Astrium (51%) and Khrunichev State Research and Pro- duction Space Center (49%) and performs commercial satellite launches with the Russian Rockot launch vehicle from modern and service-oriented spacecraft processing, launch, mission control and Customer facilities at Plesetsk Cosmodrome in Northern Russia. Rockot is the only Russian small vehicle that has benefited from any investment of this kind. Rockot builds on sufficient numbers of SS-19 stocks which are used as the two booster stages and uses as the third and upper stage the newly Khrunichev-produced Breeze-KM upper stage. The heritage of the SS-19 family comprises a total of nearly 160 flights. Next to being used by Eurockot, Rockot is also employed as a small launcher for Russian government launches. EUROCKOT Launch Services pursues a pricing policy of "no additional charges" to make launch cost transparent and better assessable: as a principle, the launch contract price agreed with the Customer will cover all expenses and services in the context of performing a launch. This also comprises duty-free importation of the spacecraft, all transport and logistical services within Russia and all launch licenses. The new issue of our User's Guide will acquaint you in detail with the capabilities of the Rockot/Breeze-KM launch system, with the management responsibilities EUROCKOT holds vis-à-vis the Customer and with the ground facilities and services we have at our disposal. The User's Guide was designed to serve you as an initial reference and I invite you to address me or my colleagues to verify your findings and with any enquiry you may have - you will be served with a quick response. With hindsight, I suggest that it is always better to contact us to draw on the knowledge and experience of my staff whose specific role it is to advise you. Sincerely,

Dr. Matthias Oehm CEO EUROCKOT Launch Services GmbH

Preface to Rockot User's Guide, EHB0003, Issue 5, Revision 0

This new issue of the Rockot User's Guide is intended to completely replace the previous version of the Rockot User's Guide Issue 4, Revision 0. EUROCKOT has now conducted 10 commercial launches and successfully injected 20 spacecraft into orbit. The Rockot Launch System described herein is therefore fully operational and integrated in the market of small launchers. All the aspects of the launch service from mission management and analyses, importation of spacecraft, launch campaign through to lift-off and spacecraft separation using a variety of flight proven adapter and separation systems have been demonstrated successfully.

Furthermore, Eurockot could increase the Rockot mass performance incrementally during the last years, mainly by mass improvements of the upper stage. In parallel the portfolio of flight qualified mechanical interfaces to the spacecraft, namely adapters and separation systems were increased during the last missions for a wide variety of Customers including both commercial and institutional and for scientific and commercial satellites since the last issue of the User's Guide.

The User's Guide has been updated to reflect all the technical and service improvements that have been obtained through our previous and existing Customers.

York Viertel Technical Director EUROCKOT Launch Services GmbH

Bremen, August 2011.

Document Change Record

Issue Number Date Revisions Approved Initial release 30.11.1994 - Muss Issue 1 20.01.1995 - Muss Issue 1/Rev 1 20.03.1995 1-1, 1-2 Muss 2-2, 2-3, Fig. 2-4 deleted 3-1, 3-2, 3-4, 3-6, 3-7, 3-8 4-6, 4-7, 4-17 5-1, 5-2, 5-3, 5-4, 5-7, 5-9 6-3 7-1, 7-2, 7-3, 7-4 8-1, 8-2, 8-3, 8-4, 8-5, 8-6, 8-8, 8- 9, 8-10, 8-12, 8-13, 8-14, 8-15, 8- 17 9-1 10-1 Issue 1/Rev 2 05.07.1995 1-2, 4-6, 4-7 P. Bamberg Issue 1/Rev 3 01.08.1995 Fig. 3.1 revised P. Bamberg 3.1.2, Fig. 3.2 – 3.4 inserted 4.1.6, Fig. 4.5 revised Issue 2/Rev 0 11.05.1998 Completely revised issue Dr. M. Kinnersley Dr. T. Miski Issue 2/Rev 1 01.06.1999 1-1, Fig. 1-4, 1-5, 1-6 changed, Dr. M. Kinnersley 2-2, Fig. 2-2, 2-3 updated, Dr. T. Miski 2-6, 2-7, 2-8, 2-9, Fig. 3-1 updated, 3-10, 3-11, Fig. 4-3 updated, 4-3, 4-4, Fig. 4-4 updated, 4-8, Table 5.1.2-1 updated, Table 5.1.3-1 updated, 5-3, Fig. 5-3, Fig. 5-6 updated, Table 5.1.7-1 updated, Table 5.4.1-1 updated, 6-1, Fig. 6-1, 6-2 updated 10-1, 10-6, Fig. 10-3 updated, 10- 7, 10-8, Fig. 10-4 updated, 10-12, Fig. 10-5 updated, 10-14, 10-15

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Issue Number Date Revisions Approved Issue 3/ Rev 0 31.1.01 Completely revised issue taking Dr. M. Kinnersley into account results of CDF and Dr. T. Miski completion of payload processing facilities. In addition, performance figures for Baikonur are included now. Issue 3/Rev 1 30.04.01 Editorial changes Dr. M. Kinnersley Dr. T. Miski Issue 4/ Rev 0 1.11.04 Completely updated issue taking Dr. M. Kinnersley into account the commercial launches of EUROCKOT. Issue 5/ Rev 0 31.7.11 Completely updated issue, in Y. Viertel particular sections 1, 3 and 4 to reflect upgrades in mass performance, mechanical interfaces and services during the last commercial launches

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Table of Contents

1 Introduction...... 1-1 1.1 Reasons for Selecting EUROCKOT...... 1-4 1.2 EUROCKOT Photograph Gallery ...... 1-4

2 Launch Vehicle Description...... 2-1 2.1 General Characteristics and Description...... 2-1 2.1.1 First Stage ...... 2-3 2.1.2 Second Stage ...... 2-4 2.1.3 Breeze-KM Upper Stage...... 2-4 2.1.4 Fairing ...... 2-6 2.1.5 Transportation and Launch Container ...... 2-7 2.2 Rockot Qualification and Flight History ...... 2-7 2.3 Revalidation of SS-19s used by EUROCKOT...... 2-10

3 General Performance Capabilities ...... 3-1 3.1 Introduction...... 3-1 3.2 Launch Azimuths and Orbit Inclinations from Plesetsk ...... 3-1 3.3 Low Earth Orbits...... 3-2 3.3.1 Payload Performance for Circular Orbits...... 3-3 3.3.2 Performance for Elliptical Orbits...... 3-3 3.3.3 Sun-Synchronous Orbits ...... 3-3 3.4 Mission Profile Description ...... 3-8 3.5 Baikonur Performance...... 3-13 3.6 Spacecraft Injection and Separation...... 3-14 3.6.1 Injection Accuracy...... 3-14 3.6.2 Spacecraft Separation ...... 3-14 3.6.2.1 Spin Stabilised Separation ...... 3-14 3.6.2.2 Three-Axis Stabilised Separation...... 3-15 3.6.2.3 Typical Multiple Satellite Deployment Scenarios ...... 3-16

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4.1.5 Payload Adapters ...... 4-12 4.1.5.1 Clamp Band Separation System Adapters...... 4-12 4.1.5.2 Mechanical Lock System Payload Adapters...... 4-17 4.2 Electrical Interfaces ...... 4-20 4.2.1 On-board Interfaces...... 4-20 4.2.1.1 Umbilical Connectors ...... 4-20 4.2.1.2 Separation Verification ...... 4-20 4.2.1.3 Interface Electrical Constraints ...... 4-20 4.2.1.4 Umbilical Harness Configuration and Specifications ...... 4-22 4.2.1.5 Matchmate / Electrical Checkout ...... 4-24 4.2.1.6 Spacecraft Electrical Interface Input Data Requirements ...... 4-24 4.2.2 Ground Electrical Interface...... 4-24 4.2.3 Payload Grounding and Bonding ...... 4-26 4.2.4 Payload Auxiliary Power Supply...... 4-28 4.2.4.1 Ground Auxiliary Power Supply ...... 4-28 4.2.4.2 In-flight Power Supply ...... 4-28 4.2.4.3 Optional Services ...... 4-28 4.2.5 Separation Ignition Command...... 4-28 4.2.6 Payload Telemetry Support ...... 4-28

5 Spacecraft Environmental Conditions...... 5-1 5.1 Mechanical Environment ...... 5-1 5.1.1 General...... 5-1 5.1.2 Quasi-Static Accelerations...... 5-2 5.1.3 Low Frequency Vibration...... 5-3 5.1.4 Acoustic Noise...... 5-4 5.1.5 Random Vibration...... 5-5 5.1.6 Shock ...... 5-5 5.1.7 Loads during Ground Operations ...... 5-7 5.1.7.1 Handling Loads During Ground Operations...... 5-8 5.1.7.2 Spacecraft Container Railway Transportation Loads ...... 5-8 5.1.7.3 Roadway Autonomous Transportation in Archangel and Plesetsk Cosmodrome...... 5-9 5.1.7.4 Loads during Transport in Upper Composite ...... 5-10 5.2 Thermal Environment ...... 5-11 5.2.1 Environmental Conditions in the Integration Facility...... 5-11 5.2.2 Pre-Launch Temperature Control within the Fairing...... 5-11 5.2.3 In-flight Temperature under the Fairing ...... 5-13 5.2.4 Aerothermal Flux at Fairing Jettisoning...... 5-14 5.2.5 Heat Impact during Coasting Phase...... 5-14 5.3 Fairing Static Pressure during the Ascent ...... 5-14 5.4 Contamination and Cleanliness...... 5-15 5.5 Electromagnetic Environment...... 5-15 5.5.1 Launch Vehicle Electro-Magnetic Emissions...... 5-15 5.5.2 Spacecraft Electro-Magnetic Emissions ...... 5-17

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6 Spacecraft Design and Verification Requirements...... 6-1 6.1 Safety Requirements...... 6-1 6.1.1 Selection of Payload Materials ...... 6-1 6.2 Design Characteristics...... 6-1 6.2.1 Mass Properties...... 6-1 6.2.2 Centre of Mass Constraints...... 6-1 6.2.3 Structural Integrity...... 6-1 6.2.3.1 Factors of Safety ...... 6-1 6.2.3.2 Dimensioning Loads...... 6-2 6.2.4 Stiffness...... 6-3 6.2.5 Overflux ...... 6-3 6.3 Spacecraft Mechanical Qualification and Acceptance Tests...... 6-3 6.3.1 Static Load Test ...... 6-3 6.3.2 Sinusoidal Vibration Test ...... 6-4 6.3.3 Random Vibration Test ...... 6-4 6.3.4 Acoustic Noise Test ...... 6-5 6.3.5 Shock Test...... 6-5 6.4 Interface Tests ...... 6-5

7 Mission Management ...... 7-1 7.1 Mission Management Overview ...... 7-1 7.2 Organisation and Responsibilities ...... 7-1 7.2.1 EUROCKOT Mission Responsibilities ...... 7-3 7.2.1.1 Mission Integration...... 7-3 7.2.1.2 Interface Design, Qualification and Verification ...... 7-4 7.2.1.2.1 Design of the Payload Adapter...... 7-4 7.2.1.2.2 Payload Adapter Qualification Test at KSRC...... 7-4 7.2.1.2.3 Fit Check of FM Spacecraft with FM Payload Adapter ...... 7-5 7.2.1.3 Configuration Control ...... 7-5 7.2.1.4 Launch Vehicle Procurement ...... 7-5 7.2.1.5 Spacecraft Preparation and Launch Operations...... 7-5 7.2.1.6 Post-Launch Activities...... 7-6 7.2.1.7 Quality /Mission Assurance...... 7-6 7.2.1.8 Safety Provisions ...... 7-6 7.2.1.9 Risk Management ...... 7-6 7.2.1.10 Technology Transfer and Security...... 7-7 7.2.2 Customer Mission Responsibilities...... 7-7 7.3 Reviews and Documentation...... 7-8 7.3.1 EUROCKOT Documents ...... 7-10 7.3.1.1 Interface Control Document ...... 7-10 7.3.1.2 Preliminary Design Review Data Package ...... 7-10 7.3.1.3 Critical Design Review Data Package ...... 7-10 7.3.1.4 Joint Operations Plan...... 7-11 7.3.1.5 Safety Reply...... 7-11 7.3.1.6 Launch Evaluation Report...... 7-11

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7.3.2 Customer Documents...... 7-11 7.4 Overall Mission Schedule...... 7-12

8 Mission Design and Mission Analysis...... 8-1 8.1 General...... 8-1 8.2 Overall Description of Spacecraft to Rockot Launch Vehicle Integration (Book 1)...... 8-3 8.3 Trajectory and Mission Sequence (Book 4)...... 8-3 8.4 Dynamics of Spacecraft Separation (Book 5)...... 8-3 8.5 Thermal Environment (Book 6, parts 1 and 2)...... 8-4 8.6 Dynamic Coupled Loads Analysis (Book 7) ...... 8-5 8.7 Spacecraft Cleanliness Control (Book 8)...... 8-6 8.8 Measurement System (Book 9)...... 8-7 8.9 Electromagnetic Compatibility Study (Book 10) ...... 8-7 8.10 Onboard Electrical Interface of Spacecraft and Launch Vehicle (Book 11) ...... 8-8 8.11 Ground Electrical Cabling, Power Supply and Interfaces with Spacecraft GSE (Book 12) ...... 8-8 8.12 Launch base operations support (Book 14 part 1, 2, 3) ...... 8-8 8.13 Reliability (Book 15, part 1) ...... 8-9 8.14 ILV Components Quality Assurance (Book 15, part 2) ...... 8-9 8.15 Social services (Book 16)...... 8-9 8.16 Communications support (Book 17) ...... 8-9 8.17 Security (Book 18)...... 8-9 8.18 Transportation from Port-of-Entry to the Launch Site Facilities (Book 19) ...... 8-10

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9.3.7 Guidelines for Safety Analyses...... 9-4 9.3.7.1 Overall Assessment of Risk and Severity ...... 9-4 9.3.7.2 Probability of Hazards...... 9-5 9.3.7.3 Prevention of Hazards...... 9-5 9.3.7.4 Reference Documents ...... 9-5 9.4 Non-compliance with Safety Requirements / Waivers...... 9-6 9.5 Summary ...... 9-6

10 Plesetsk Cosmodrome...... 10-1 10.1 General Description...... 10-1 10.1.1 Climatic Conditions...... 10-3 10.2 Logistics ...... 10-3 10.2.1 Spacecraft and Hardware Transport...... 10-3 10.2.2 Transportation Requirements ...... 10-4 10.2.3 Transport Environments ...... 10-5 10.2.3.1 Environmentally Controlled Transport of Spacecraft during Launch Campaign ...... 10-5 10.2.4 Spacecraft Team Transport to Plesetsk Cosmodrome...... 10-5 10.2.4.1 Charter Aircraft ...... 10-5 10.2.4.2 Scheduled Flights...... 10-6 10.2.4.3 Rail Transfer to Plesetsk Cosmodrome ...... 10-6 10.2.5 Customer Team Transport at the Launch Site ...... 10-6 10.3 Communications...... 10-6 10.3.1 Phone Lines...... 10-7 10.3.2 Data Lines...... 10-7 10.3.3 Mobile Radios...... 10-8 10.3.4 CCTV...... 10-8 10.3.5 Entertainment TV...... 10-9 10.4 Ground Facilities...... 10-9 10.4.1 The Integration Facility MIK...... 10-9 10.4.1.1 General Hall ...... 10-12 10.4.1.2 Clean Room Bay ...... 10-12 10.4.1.2.1 Airlock ...... 10-12 10.4.1.2.2 Upper Composite Integration Area ...... 10-12 10.4.1.2.3 Spacecraft Processing Area ...... 10-14 10.4.1.3 EGSE Rooms...... 10-15 10.4.1.4 Customer's Office Area ...... 10-16 10.4.1.5 Handling and Hoisting Equipment in MIK ...... 10-16 10.4.1.6 Power Supply of MIK...... 10-16 10.4.2 The Rockot Launch Complex ...... 10-19 10.4.2.1 Launch Pad ...... 10-20 10.4.2.2 Undertable Rooms ...... 10-21 10.4.2.3 Power Supply of Launch Complex...... 10-21 10.4.2.4 Air Conditioning of the Spacecraft at the Launch Pad ...... 10-25 10.4.3 The Mission Control Centre...... 10-25

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10.5 Launch Campaign ...... 10-26 10.5.1 Responsibilities and Operational Organisation...... 10-26 10.5.2 Planning ...... 10-26 10.5.3 Procedures and Logbook of Works ...... 10-27 10.5.4 Training / Briefings...... 10-27 10.5.5 Security and Access Control ...... 10-27 10.5.6 Safety ...... 10-28 10.5.7 Launch Campaign Operations...... 10-28 10.5.7.1 Launch Vehicle Operations in MIK and at Launch Complex...... 10-30 10.5.7.2 Spacecraft Operations...... 10-30 10.5.7.3 Combined Operations in MIK...... 10-30 10.5.7.4 Combined Operations at the Launch Pad...... 10-31 10.5.8 Launch Day...... 10-32 10.5.9 Abort Re-Cycle/ Return-to-Base Operation...... 10-33 10.6 Accommodation and Leisure Activities...... 10-34 10.7 Medical Care...... 10-34

11 Baikonur Cosmodrome...... 11-1

12 Items to be Delivered by the Customer ...... 12-1 12.1 General Documents ...... 12-2 12.1.1 Interface Requirements Document...... 12-2 12.1.2 Orbit Tracking Operation Report...... 12-3 12.2 Input to Mission Design and Mission Analysis...... 12-3 12.2.1 Response to Questionnaire...... 12-3 12.2.2 Spacecraft Dynamic Model...... 12-5 12.2.3 Thermal Model ...... 12-5 12.3 Safety ...... 12-6 12.4 Payload Environmental Test Documents...... 12-6 12.5 Operations Documents for Spacecraft...... 12-6 12.6 Contractual / Higher Level Documents...... 12-7 12.7 Models, GSE...... 12-7 12.8 Hardware, Software and Document Time Schedule ...... 12-8

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List of Figures

Figure 1-1 Shareholders in the EUROCKOT joint venture ...... 1-3 Figure 1-2 EUROCKOT roles and responsibilities...... 1-3 Figure 1-3 Rockot launch from Plesetsk Cosmodrome (GRACE, 2002)...... 1-5 Figure 1-4 Arrival of spacecraft and support equipment at Archangel Talagi airport, the Russian port of entry...... 1-6 Figure 1-5 Transfer of the spacecraft container to Archangelskaya train station for rail transport to the Plesetsk Cosmodrome...... 1-7 Figure 1-6 Loading of spacecraft and support equipment at Archangelskaya train station...... 1-7 Figure 1-7 Thermally controlled railroad car for transportation of specific GSE...... 1-8 Figure 1-8 Arrival of the spacecraft container in the general hall of the integration and test facility MIK...... 1-8 Figure 1-9 Optional spacecraft air transport to the airfield of the Plesetsk Cosmodrome, Ilyushin IL-76...... 1-9 Figure 1-10 EUROCKOT test and integration facility (MIK) with Rockot booster in the foreground...... 1-9 Figure 1-11 Rockot booster delivery to the launch pad...... 1-10 Figure 1-12 Booster installation at the launch pad...... 1-10 Figure 1-13 Unpacking of GSE containers in EUROCKOT clean rooms...... 1-11 Figure 1-14 Spacecraft processing in the EUROCKOT payload processing facility (KOMPSAT-2, 2006)...... 1-11 Figure 1-15 Spacecraft processing in the EUROCKOT payload processing facility (SERVIS-1, 2003)...... 1-12 Figure 1-16 Spacecraft installation on the multi-satellite dispenser (GRACE, 2002)...... 1-12 Figure 1-17 Control centre of EUROCKOT closed circuit television system for Customer usage...... 1-13 Figure 1-18 Spacecraft hydrazine fuelling on EUROCKOT fuelling platform...... 1-14 Figure 1-19 Spacecraft mating with the adapter system using an EADS CASA clamp band...... 1-14 Figure 1-20 Transfer of integrated spacecraft and adapter for mating operations with the Breeze-KM upper stage (SERVIS-1, 2003)...... 1-15 Figure 1-21 Mating with the Breeze-KM upper stage (SERVIS-1, 2003)...... 1-16 Figure 1-22 Mating of GOCE spacecraft (2009)...... 1-17 Figure 1-23 Final encapsulation operations (GRACE, 2002)...... 1-18 Figure 1-24 Final encapsulation operations (SERVIS-1, 2003)...... 1-19 Figure 1-25 Completion of payload encapsulation (GRACE, 2002)...... 1-20 Figure 1-26 Rockot upper composite with thermal conditioning unit rail car protected by thermal cover for transfer to the launch pad...... 1-21 Figure 1-27 Rockot upper composite on its way to the launch pad...... 1-22 Figure 1-28 Hoisting of upper composite at the Rockot service tower...... 1-23 Figure 1-29 Stacking of upper composite on the Rockot booster...... 1-24 Figure 1-30 Rockot ready for launch with the mobile service tower retracted...... 1-25 Figure 1-31 Rockot launch (Iridium, 2002)...... 1-26 Figure 1-32 EUROCKOT remote mission control centre (MCC)...... 1-26

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Figure 1-33 Display of live telemetry in the MCC...... 1-27 Figure 1-34 Display of trajectory with actual coordinates in the MCC...... 1-28 Figure 1-35 Meeting rooms and dining hall at the Rockot Hotel...... 1-29 Figure 1-36 Plentiful culture and nature invite to relax outside the cosmodrome...... 1-30 Figure 1-37 Arrival of a EUROCKOT charter flight at the cosmodrome's airfield...... 1-31 Figure 1-38 Breeze-KM upper stage production at KSRC, Moscow...... 1-31 Figure 1-39 Aft view of the Breeze-KM propulsion compartment...... 1-32 Figure 1-40 Spacecraft separation shock test at the Customer's facilities (GOCE, 2008)...... 1-32 Figure 1-41 Typical spacecraft separation test using simulators at KSRC facilities, Moscow...... 1-33 Figure 2-1 Rockot launch vehicle configuration...... 2-2 Figure 2-2 Upper composite with payload...... 2-5 Figure 3-1 Rockot/Breeze-KM payload performance for circular orbits...... 3-4 Figure 3-2 Rockot/Breeze-KM payload performance for elliptical orbits at 63.2° inclination...... 3-5 Figure 3-3 Rockot/Breeze-KM payload performance for elliptical orbits at 72.0° inclination...... 3-6 Figure 3-4 Rockot/Breeze-KM payload performance for elliptical orbits at 82.5° inclination...... 3-7 Figure 3-5 Cycling orientation of Breeze-KM relative to the sun during coast flight maintained for one hour (a), and for half an hour (b)...... 3-10 Figure 3-6 Typical Rockot/Breeze-KM ascent (650 km circular, 86.583° inclination). (a) Trajectory. (b) Acceleration timeline...... 3-17 Figure 3-7 Typical Rockot/Breeze-KM ascent (1000 km circular, 99.52° inclination). (a) Trajectory. (b) Acceleration timeline...... 3-19 Figure 3-8 Typical Rockot/Breeze-KM ascent (320 km perigee, 820 km apogee at 96.8° inclination and SSO 820 km, at 98.7° inclination). (a) Trajectory. (b) Acceleration timeline...... 3-21 Figure 3-9 Sun-synchronous orbit injection scheme...... 3-22 Figure 3-10 Example of sequential deployment of three spacecraft...... 3-25 Figure 3-11 Example of simultaneous deployment of six spacecraft...... 3-26 Figure 4-1 Multiple payload accommodation for MOM without main payload...... 4-2 Figure 4-2 Integrated MOM payload with main payload simulator...... 4-2 Figure 4-3 Rockot maximum usable payload envelope, all dimensions given in mm.....4-4 Figure 4-4 CASA CRSS Clamp Ring Separation System installation with KSRC pyrolock...... 4-6 Figure 4-5 Stay-out zones (hatched) for the standard interface diameter of 937 mm, all dimensions given in mm...... 4-7 Figure 4-6 Stay-out zones (hatched) for the standard interface diameter of 1194 mm, all dimensions given in mm...... 4-8 Figure 4-7 CASA LPSS Launcher Payload Separation System, (a) general view, (b) stay-out zones (hatched) for the standard interface diameter of 1194 mm, all dimensions given in mm...... 4-9 Figure 4-8 Cut-away detail of the Mechanical Lock System...... 4-10 Figure 4-9 Minisatellite separation system...... 4-11 Figure 4-10 CASA 937 SRF clamp band cylindrical payload adapter. (a) Side view. (b) Launch vehicle interface ring detail...... 4-14

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Figure 4-11 CASA CRSS 1194 SRF clamp band conical payload adapter system (Kompsat-2). (a) Side view. (b) Launch vehicle interface ring, detail...... 4-15 Figure 4-12 Elongated conical payload adapter for the CASA CRSS 937 SRF clamp band...... 4-16 Figure 4-13 Example of base-mounted multiple satellite dispenser system for two spacecraft (SMOS/Proba-2)...... 4-16 Figure 4-14 Example of an MLS adapter system for single satellite accommodation. ...4-18 Figure 4-15 Example of side-mounted multiple satellite dispenser system for two spacecraft (GRACE). Please note that only one spacecraft shown in this photo...... 4-19 Figure 4-16 Example of side-mounted multiple satellite dispenser system for three spacecraft (SWARM)...... 4-19 Figure 4-17 Typical example of an umbilical connector bracket used in combination with a 1194 mm clamp band...... 4-21 Figure 4-18 Umbilical connector OSRS50BATV...... 4-21 Figure 4-19 Rockot umbilical harness diagram...... 4-23 Figure 4-20 Launch site ground wiring diagram...... 4-25 Figure 5-1 Launch vehicle and payload coordinate system...... 5-1 Figure 5-2 Typical axial payload interface acceleration...... 5-2 Figure 5-3 Typical radial payload interface acceleration...... 5-3 Figure 5-4 Low frequency vibration environment at the separation plane...... 5-4 Figure 5-5 Composite noise spectrum under the Rockot payload fairing during flight given in 1/3 octave band referenced to 2 x 10-5 Pa RMS...... 5-4 Figure 5-6 Typical shock environment for diverse separation systems...... 5-6 Figure 5-7 Transportation of the spacecraft in the spacecraft container...... 5-7 Figure 5-8 Spacecraft transportation as part of the upper composite...... 5-7 Figure 5-9 Random vibration power spectrum acting at spacecraft container base during autonomous railway transportation...... 5-8 Figure 5-10 Random vibration spectrum for autonomous roadway transportation...... 5-9 Figure 5-11 Random vibration power spectrum acting on the spacecraft during upper composite transportation...... 5-10 Figure 5-12 Thermal conditioning of the upper composite during transportation...... 5-12 Figure 5-13 Thermal conditioning of the upper composite at the launch pad...... 5-12 Figure 5-14 Variation of fairing inner static pressure during ascent...... 5-15 Figure 5-15 Field strength of launch vehicle and cosmodrome electromagnetic emissions in the adapter plane without payload fairing...... 5-16 Figure 5-16 Allowable spacecraft emissions at Plesetsk Cosmodrome...... 5-17 Figure 6-1 Typical limit load factors for initial dimensioning of secondary structures and equipment brackets...... 6-2 Figure 7-1 Industrial organisation of EUROCKOT and its major subcontractors...... 7-2 Figure 7-2 EUROCKOT mission management organisation...... 7-3 Figure 9-1 Safety submission phases...... 9-2 Figure 10-1 Geographical location of the Plesetsk Cosmodrome...... 10-2 Figure 10-2 Plesetsk Cosmodrome layout...... 10-3 Figure 10-3 Communications links...... 10-9 Figure 10-4 Layout of the integration facility MIK...... 10-14 Figure 10-5 Assembly stand with platforms in upper composite integration room...... 10-16 Figure 10-6 Scheme of removable fuelling platform...... 10-19

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Figure 10-7 Customer office area in MIK...... 10-22 Figure 10-8 Launch complex for Rockot...... 10-25 Figure 10-9 Rockot launch pad...... 10-28 Figure 10-10 Undertable Room 7 for Customer's use...... 10-28 Figure 10-11 Schedule of operations...... 10-34 Figure 10-12 Flow of combined operations in MIK...... 10-36 Figure 10-13 Launch pad operations...... 10-37 Figure 12-1 Interface Requirements Document table of contents...... 12-2 Figure 12-2 Spacecraft Operations Plan table of contents...... 12-7 Figure 12-3 Dates of the Customer's documents, software and hardware supply...... 12-9

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List of Tables

Table 2-1 Main characteristics of the Rockot/Breeze-KM launch vehicle...... 2-1 Table 2-2 First stage main engine characteristics...... 2-3 Table 2-3 Second stage main engine characteristics...... 2-4 Table 2-4 Second stage vernier thrusters characteristics...... 2-4 Table 2-5 Breeze-KM upper stage main engine characteristics...... 2-6 Table 2-6 Breeze-KM vernier engine characteristics...... 2-6 Table 2-7 Breeze-KM AOCS engine characteristics...... 2-6 Table 2-8 Estimated Breeze-KM upper stage mass...... 2-6 Table 2-9 EUROCKOT launch record with Rockot/Breeze-KM...... 2-9 Table 3-1 Allowable launch azimuths that can be served from Plesetsk...... 3-1 Table 3-2 Orbital injection accuracy...... 3-14 Table 4-1 Pin allocation of umbilical connectors...... 4-23 Table 4-2 Transit wire configurations...... 4-24 Table 4-3 Bonding/grounding schematic drawing; example of a dispenser configuration...... 4-27 Table 4-4 Operational parameters of the telemetry system TA1...... 4-29 Table 4-5 Operational parameters of the telemetry system TA2...... 4-29 Table 5-1 Quasi-static loads experienced by an average payload during ascent...... 5-3 Table 5-2 Low frequency vibration environment...... 5-3 Table 5-3 Composite noise spectrum under the fairing during flight...... 5-5 Table 5-4 Typical shock environment for diverse separation systems...... 5-6 Table 5-5 Maximum handling loads...... 5-8 Table 5-6 Shock loads during autonomous railway transportation...... 5-8 Table 5-7 Random vibration power spectrum acting at spacecraft container base during autonomous railway transportation...... 5-8 Table 5-8 Quasi-static accelerations for autonomous roadway transportation at wind velocities ≤ 20 m/s...... 5-9 Table 5-9 Random vibration power spectrum for autonomous roadway transportation...... 5-9 Table 5-10 Accelerations during upper composite transportation at wind velocities ≤ 20 m/s...... 5-10 Table 5-11 Random vibration power spectrum acting on the spacecraft during upper composite transportation...... 5-10 Table 5-12 Air conditioning systems performance data...... 5-13 Table 5-13 Measured parameters under the payload fairing...... 5-13 Table 5-14 Parameters of the launch vehicle transmitters...... 5-16 Table 5-15 Restriction on radio frequency use by the spacecraft during launch...... 5-17 Table 6-1 Sinusoidal vibration loads...... 6-4 Table 6-2 Typical mechanical verification test matrix for qualification (Q) and acceptance (A) of the spacecraft...... 6-4 Table 6-3 Acoustic noise spectrum...... 6-5 Table 7-1 Typical launch services reviews and meetings...... 7-9 Table 7-2 Typical ICD structure...... 7-10 Table 7-3 Documents to be supplied by the Customer...... 7-12

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Table 7-4 Typical mission schedule...... 7-13 Table 8-1 Data package structure...... 8-2 Table 8-2 Data package book content...... 8-2 Table 10-1 Characteristics of airports for shipping of spacecraft containers, related GSE and personnel transfer...... 10-5 Table 10-2 Specification of 208/120 V 60 Hz AC processing facility power supply system...... 10-23 Table 10-3 Specification of 380/220 V 50 Hz AC processing facility uninterrupted power supply system...... 10-24 Table 12-1 Summary of documents to be supplied by the customer...... 12-1

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Acronyms and Abbreviations

AC Alternating Current AOCS Attitude and Orbit Control System

CCTV Closed-circuit Television CDF Commercial Demonstration Flight CDR Critical Design Review CEO Chief Executive Officer CIS Commonwealth of Independent States CLA Coupled Loads Analysis CoG Centre of Gravity COP Combined Operations Plan

DC Direct Current DPA Destructive Physical Analysis DT Direct Transmission Mode (Telemetry) DTSA Defense Technology Security Administration DQR Design Qualification Review

EGSE Electrical Ground Support Equipment EOGS End of Gyro Setting (= Mission Zero Time) EMC Electromagnetic Compatibility

FMAD Final Mission Analysis Documentation FMAR Final Mission Analysis Review FS Factor of Safety

GMF Ground Measurement Facility GMI Ground Measurement Infrastructure GN2 Gaseous Nitrogen GPS Global Positioning System (US), GLONASS Global Navigation Space System (CIS) GSE Ground Support Equipment

ICBM Intercontinental Ballistic Missile ICD Interface Control Document ILV Integrated Launch Vehicle IRD Interface Requirements Document

JOP Joint Operations Plan

KSRC Khrunichev State Research and Production Space Center

LCR Launch Control Room LER Launch Evaluation Report LOS Launch Operation Schedule LRD Launch Requirements Document LRR Launch Readiness Review

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LSM Launch Service Manager LV Launch Vehicle LVM Launch Vehicle Manager MA Mission Assurance MAR Mission Analysis Review MCC Mission Control Centre MGSE Mechanical Ground Support Equipment MIK Spacecraft Integration Facility for ROCKOT in Plesetsk MLS Mechanical Lock System MM Mission Manager MoI Moment of Inertia

N2O4 Nitrogen Tetroxide (Oxidizer)

O(A)SPL Overall Sound Pressure Level

PDR Preliminary Design Review PLF Payload Fairing PMAD Preliminary Mission Analysis Documentation PMAR Preliminary Mission Analysis Review PoE Port of Entry PSD Power Spectral Density

QA Quality Assurance

RAAN Right Ascension of Ascending Node REC Data Record Mode (Telemetry) REP Data Replay Mode (Telemetry) RF Radio Frequency RMS Root Mean Square RPM Revolutions per Minute

SC Spacecraft SDR Systems Design Review SMD Spacecraft Mission Director SOP Spacecraft Operations Plan SOTP Spacecraft Operations Test Procedure SPPA Single Pyro Released Point Attachment System SSO Solar Synchronous Orbit

TA1 Rockot Low Rate Telemetry Device TA2 Rockot High Rate Telemetry Device TAA Technical Assistance Agreement TIM Technical Interchange Meeting TLC Transport and Launch Container

UDMH Unsymmetrical Dimethyl Hydrazine (Fuel)

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Chapter 1 Introduction

Table of Contents

1 Introduction ...... 1-1 1.1 Reasons for Selecting EUROCKOT...... 1-4 1.2 EUROCKOT Photograph Gallery ...... 1-4

List of Figures

Figure 1-1 Shareholders in the EUROCKOT joint venture ...... 1-2 Figure 1-2 EUROCKOT roles and responsibilities...... 1-3 Figure 1-3 Rockot launch from Plesetsk Cosmodrome (GRACE, 2002)...... 1-5 Figure 1-4 Arrival of spacecraft and support equipment at Archangel Talagi airport, the Russian port of entry...... 1-6 Figure 1-5 Transfer of the spacecraft container to Archangel train station for rail transport to the Plesetsk Cosmodrome...... 1-7 Figure 1-6 Loading of spacecraft and support equipment at Archangel train station. ....1-7 Figure 1-7 Thermally controlled railroad car for transportation of specific GSE...... 1-8 Figure 1-8 Arrival of the spacecraft container in the general hall of the integration and test facility MIK...... 1-8 Figure 1-9 Optional spacecraft air transport to the airfield of the Plesetsk Cosmodrome using Ilyushin IL-76...... 1-9 Figure 1-10 EUROCKOT test and integration facility (MIK) with Rockot booster in the foreground...... 1-9 Figure 1-11 Rockot booster delivery to the launch pad...... 1-10 Figure 1-12 Booster installation at the launch pad...... 1-10 Figure 1-13 Unpacking of GSE containers in EUROCKOT clean rooms...... 1-11 Figure 1-14 Spacecraft processing in the EUROCKOT payload processing facility (KOMPSAT-2, 2006)...... 1-11 Figure 1-15 Spacecraft processing in the EUROCKOT payload processing facility (SERVIS-1, 2003)...... 1-12 Figure 1-16 Spacecraft installation on the multi-satellite dispenser (GRACE, 2002)...... 1-12 Figure 1-17 Controls of EUROCKOT Closed Circuit Television (CCTV) system for Customer usage...... 1-13 Figure 1-18 Spacecraft hydrazine fuelling on EUROCKOT fuelling platform...... 1-14 Figure 1-19 Spacecraft mating with the adapter system using an EADS CASA clamp band...... 1-14 Figure 1-20 Transfer of integrated spacecraft and adapter for mating operations with the Breeze-KM upper stage (SERVIS-1, 2003)...... 1-15 Figure 1-21 Mating with the Breeze-KM upper stage (SERVIS-1, 2003)...... 1-16 Figure 1-22 Mating of GOCE spacecraft (2009)...... 1-17 Figure 1-23 Final encapsulation operations (GRACE, 2002)...... 1-18 Figure 1-24 Final encapsulation operations (SERVIS-1, 2003)...... 1-19

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Figure 1-25 Completion of payload encapsulation (GRACE, 2002)...... 1-20 Figure 1-26 Rockot upper composite with thermal conditioning unit rail car protected by thermal cover for transfer to the launch pad...... 1-21 Figure 1-27 Rockot upper composite on its way to the launch pad...... 1-22 Figure 1-28 Hoisting of upper composite at the Rockot service tower...... 1-23 Figure 1-29 Stacking of upper composite onto the Rockot booster...... 1-24 Figure 1-30 Rockot ready for launch with the mobile service tower retracted...... 1-25 Figure 1-31 Rockot launch (Iridium, 2002)...... 1-26 Figure 1-32 EUROCKOT remote Mission Control Centre (MCC)...... 1-26 Figure 1-33 Display of live telemetry in the MCC...... 1-27 Figure 1-34 Display of trajectory with actual coordinates in the MCC...... 1-28 Figure 1-35 Meeting rooms and dining room at the Rockot Hotel...... 1-29 Figure 1-36 Plentiful culture and nature invite to relax outside the cosmodrome...... 1-30 Figure 1-37 Arrival of a EUROCKOT charter flight at the cosmodrome's airfield...... 1-31 Figure 1-38 Breeze-KM upper stage production at KSRC, Moscow...... 1-31 Figure 1-39 Aft view of the Breeze-KM propulsion compartment...... 1-32 Figure 1-40 Spacecraft separation shock test at the Customer's facilities (GOCE, 2008)...... 1-32 Figure 1-41 Typical spacecraft separation test using simulators at KSRC facilities, Moscow...... 1-33

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1 Introduction Unmanned Space Experiment Free Flyer (USEF).

Today, the Rockot/Breeze-KM launch vehicle This User’s Guide is provided by through its successful commercial launches, EUROCKOT Launch Services to familiarise is now a firmly established and operational potential customers with commercial launch system offering a complete all inclusive trans- services using the operational Russian portation service to low earth orbit and be- Rockot launch vehicle. The guide provides yond. background information on the Rockot launch vehicle as well as information on the EUROCKOT's previous missions and whole range of launch service activities. This launch campaigns have now fully demon- includes the heritage and previous flight his- strated the major capabilities of the Rockot tory of Rockot, a technical overview of the system and associated launch services. complete launcher, payload performance, This includes all the aspects of the incom- payload accommodation and interfaces, ing transportation logistics for spacecraft launch environment, payload dimensioning importation, launch campaigns with sophis- and verification requirements, safety, mission ticated spacecraft, different types of trajec- management, mission analysis, the launch tories such as polar and sun-synchronous site and operations. orbits, multiple re-start capability of the upper stage for plane changes and orbit rais- This current issue of the User's Guide has ing, injection of spacecraft into different been extensively updated to take into ac- orbits in one mission and more. count the experience gained since the last edition. Following EUROCKOT's successful With launches being offered from Europe's maiden launch in 2000 with the Commercial closest orbital launch site together with com- Demonstration Flight many commercial petitive terms and conditions, EUROCKOT is payloads and missions have been suc- well placed to serve the world's small satel- cessfully launched by EUROCKOT. lite community in the years to come. As well EUROCKOT's customers include many of as LEO launches for scientific, earth obser- the world's major space agencies and sat- vation and commercial telephony/ messag- ellite manufacturers. ing payloads and constellations, innovative solutions are offered for customers requiring The Rockot family in all its variants has now launches to higher orbits and interplanetary completed over 160 launches. The commer- missions, enabling such missions to be ac- cial version of the Rockot launch vehicle is complished reliably and at an affordable equipped with the Breeze-KM upper stage price. Furthermore through incremental (Rockot/Breeze-KM) and is operated by product improvements which preserve the EUROCKOT. Up to now EUROCKOT has historical reliability of the launcher, orbited a total of more than 30 spacecraft EUROCKOT aims to stay as the leader in including two GRACE spacecraft from a joint this market segment. DLR/NASA project, two Iridium spacecraft and Japan's SERVIS-1 from the institute of Thanks to the use of existing SS-19 assets as the basis for the launch vehicle,

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EUROCKOT is able to offer the Rockot industrial partner for all legal aspects. launch system under highly competitive EUROCKOT is a company established un- terms and conditions. In addition, der German law and offers all legal safe- EUROCKOT offers a 24-month cycle be- guards provided by a western company. tween contract signature and launch as a As a partner and constituent company of baseline. Finally, the use of the modern re- EUROCKOT, Khrunichev Space Center ignitable Breeze-KM upper stage allows (KSRC) of Russia provides the launch vehi- high injection accuracy and complex ma- cle and the launch services under sub- noeuvres to be achieved. contract to EUROCKOT (see Figure 1-2). The Rockot launch vehicle is marketed and The Rockot programme is guaranteed by a operated under the aegis of the German- Russian Government decree which author- Russian joint venture company “EUROCKOT ises the EUROCKOT joint venture to use Launch Services” which was founded in these missiles and the dedicated launch March 1995 to exclusively market and per- pad and integration facilities in Plesetsk. form launch services with this vehicle. Rockot uses existing SS-19 ICBM assets Astrium GmbH of Germany holds 51% of the retired under the Strategic Arms Reduction capital, with Khrunichev State Research and Treaty for its first and second stages. A Production Space Center (KSRC) of the Rus- third stage, Breeze-KM, is added to sian Federation holding the remaining 49% as achieve orbital injection. shown in Figure 1-1. A sufficient number of SS-19 booster units EUROCKOT is the interface to the cus- for EUROCKOT use is at the disposal of tomer. It is responsible for all commercial KSRC. If the need arises many more than activities, launch contract conclusion and 100 SS-19s are available at Russian De- launch implementation as the single prime fence Organisations for potential contractor for the customer and as the sole EUROCKOT purposes.

KHRUNICHEV State Research and Production Space Center

51% 49%

Figure 1-1 Shareholders in the EUROCKOT joint venture

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KHRUNICHEV State Research and Production Space Center

Prime contractor

Launch contract Launch vehicle production Technical and management support upon request Mission management including Mission analysis, payload analytical and operational coordination physical integration Financial backing

Launch vehicle and services Launch operations at Plesetsk Logistic support procurement Cosmodrome

Technical and hardware Flight evaluation management Safety Insurance Quality Financial arrangements Mission assurance (reliability)

Figure 1-2 EUROCKOT roles and responsibilities.

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1.1 Reasons for Selecting • Strong, well-established parent compa- EUROCKOT nies, namely Astrium GmbH and KSRC

EUROCKOT offers the following unique • Rapid launch rates and launch cam- benefits and discriminators as a launch paigns service provider: • State-of-the-art dedicated launch op- • A highly reliable, operational launch erations facilities including pad and in- system available today. tegration facilities in Plesetsk

• A complete and proven launch ser- • The Rockot launcher is part of the vices package with highly competitive Russian long-term space programme terms and prices. backed by guarantees from the Rus- sian and German Governments • A customer-oriented, comprehensive range of services, including logistics 1.2 EUROCKOT Photograph for the spacecraft and associated Gallery equipment.

• Experienced EUROCKOT and subcon- Photographs of the Rockot launch vehicle tractor teams assuring a high quality system including the launch vehicle, facili- service. ties and factory are shown on the following pages (Figure 1-3 to Figure 1-41). • High accuracy of spacecraft injection and re-ignition capability with the modern Breeze-KM upper stage for flexibility in mission design.

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Figure 1-3 Rockot launch from Plesetsk Cosmodrome (GRACE, 2002).

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Figure 1-4 Arrival of spacecraft and support equipment at Archangel Talagi airport, the Russian port of entry.

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Figure 1-5 Transfer of the spacecraft container to Archangel train station for rail transport to the Ple- setsk Cosmodrome.

Figure 1-6 Loading of spacecraft and support equipment at Archangel train station.

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Figure 1-7 Thermally controlled railroad car for transportation of specific GSE.

Figure 1-8 Arrival of the spacecraft container in the general hall of the integration and test facility MIK.

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Figure 1-9 Optional spacecraft air transport to the airfield of the Plesetsk Cosmodrome using Ilyushin IL-76.

Figure 1-10 EUROCKOT test and integration facility (MIK) with Rockot booster in the foreground.

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Figure 1-11 Rockot booster delivery to the launch pad.

Figure 1-12 Booster installation at the launch pad.

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Figure 1-13 Unpacking of GSE containers in EUROCKOT clean rooms.

Figure 1-14 Spacecraft processing in the EUROCKOT payload processing facility (KOMPSAT-2, 2006).

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Figure 1-15 Spacecraft processing in the EUROCKOT payload processing facility (SERVIS-1, 2003).

Figure 1-16 Spacecraft installation on the multi-satellite dispenser (GRACE, 2002).

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Figure 1-17 Controls of EUROCKOT Closed Circuit Television (CCTV) system for Customer usage.

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Figure 1-18 Spacecraft hydrazine fuelling on EUROCKOT fuelling platform.

Figure 1-19 Spacecraft mating with the adapter system using an EADS CASA clamp band.

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Figure 1-20 Transfer of integrated spacecraft and adapter for mating operations with the Breeze-KM upper stage (SERVIS-1, 2003).

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Figure 1-21 Mating with the Breeze-KM upper stage (SERVIS-1, 2003).

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Figure 1-22 Mating of GOCE spacecraft (2009).

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Figure 1-23 Final encapsulation operations (GRACE, 2002).

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Figure 1-24 Final encapsulation operations (SERVIS-1, 2003).

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Figure 1-25 Completion of payload encapsulation (GRACE, 2002).

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Figure 1-26 Rockot upper composite with thermal conditioning unit rail car protected by thermal cover for transfer to the launch pad.

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Figure 1-27 Rockot upper composite on its way to the launch pad.

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Figure 1-28 Hoisting of upper composite at the Rockot service tower.

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Figure 1-29 Stacking of upper composite onto the Rockot booster.

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Figure 1-30 Rockot ready for launch with the mobile service tower retracted.

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Figure 1-31 Rockot launch (Iridium, 2002).

Figure 1-32 EUROCKOT remote Mission Control Centre (MCC).

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Figure 1-33 Display of live telemetry in the MCC.

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Figure 1-34 Display of trajectory with actual coordinates in the MCC.

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Figure 1-35 Meeting rooms and dining room at the Rockot Hotel.

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Figure 1-36 Plentiful culture and nature invite to relax outside the cosmodrome.

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Figure 1-37 Arrival of a EUROCKOT charter flight at the cosmodrome's airfield.

Figure 1-38 Breeze-KM upper stage production at KSRC, Moscow.

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Figure 1-39 Aft view of the Breeze-KM propulsion compartment.

Figure 1-40 Spacecraft separation shock test at the Customer's facilities (GOCE, 2008).

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Figure 1-41 Typical spacecraft separation test using simulators at KSRC facilities, Moscow.

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Chapter 2 Launch Vehicle Description

Table of Contents

2 Launch Vehicle Description ...... 2-1 2.1 General Characteristics and Description...... 2-1 2.1.1 First Stage ...... 2-3 2.1.2 Second Stage ...... 2-4 2.1.3 Breeze-KM Upper Stage...... 2-4 2.1.4 Fairing...... 2-6 2.1.5 Transport and Launch Container ...... 2-7 2.2 Rockot Qualification and Flight History ...... 2-7 2.3 Revalidation of SS-19s used by EUROCKOT...... 2-10

List of Figures

Figure 2-1 Rockot launch vehicle configuration...... 2-2 Figure 2-2 Upper composite with payload...... 2-5

List of Tables

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2 Launch Vehicle Figure 1-3 depicts a Rockot launch from Description Plesetsk. From Plesetsk Cosmodrome, Rockot is launched above ground from a conventional launch pad. However, it is still 2.1 General Characteristics and launched from the same transport and Description launch Container (TLC) that is used for the silo launches. This is to retain the com- Rockot/Breeze-KM is a fully operational, monality and heritage of the previous mis- three stage, liquid propellant Russian sile launches. launch vehicle being offered commercially by EUROCKOT Launch Services for The Rockot vehicle offered by EUROCKOT launches into low earth orbit. EUROCKOT, is a commercialised version of the basic a German-Russian joint venture company, Rockot vehicle launched three times from was formed specifically to offer this vehicle Baikonur. This commercial version, the commercially. Rockot launch vehicle with the Breeze-KM upper stage, is the only version to be of- The Rockot launch vehicle uses the SS- fered by EUROCKOT and is fully described 19/RS-18 Stiletto ICBM for its first two in the following sections of this chapter. stages. The SS-19, which was originally developed as the Russian UR-100N ICBM Table 2-1 provides an overview of the main series, was designed between 1964 and characteristics of the launcher. 1975. Over 360 SS-19 ICBMs were manufactured during the 70s and 80s. A photograph of the SS-19 ICBM being transported to EUROCKOT's launch pad at Ple- setsk Cosmodrome can be seen in Figure Characteristic Value 1-11. The Breeze-KM upper stage uses a Lift-off mass 107 tons re-startable storable liquid propellant en- Number of stages 3 gine that has been used in many other So- Fuel N2O4 / UDMH for all 3 stages viet space projects. Length 29.15 m External diameter 2.50 m (Payload fairing 2.5 x 2.62 m) Maximum payload 2140 kg into 200 km circular performance orbit inclined at 63.2°

Table 2-1 Main characteristics of the Rockot/ Breeze-KM launch vehicle.

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470 30° 2620 A-A II 1260 22°

1294 2500 III I 18.55° 6735

A-A IV FairingPayload Static fairing Envelope static envelope 3711 A-AA-A

AdapterPayload adapter ng Payload fairing EquimentEquipment Compartment compartment / Dispenser / dispenser UDMHUDMH Tank tank

NO24 Tank N2O4 tank EngineEngine Compartment compartment and and Tanks tanks ge Upper Composite Installation Plane Breeze-KM upper stage Upper composite installation plane

NO24 Tank N2O4 tank

UDMHUDMH Tank tank Verniers ter Vernier thrusters (RD-0233) 2nd stage booster 2nd2nd Stage stage Main main Engine engine (RD0235) (RD-0235)

29150

Ø 2500

NO24 Tank N2O4 tank

UDMH Tank tank

Main Characteristics of the Rockot Launch Vehic

Characteristic Value Lift-off Mass (metric tons) 107 Number of Stages 3 Fuel N2O4 / UDMH fo Overall Length (m) 29 External Diameter (m) 2.5 (PLF 2.6) Max Payload Performance 1950 kg into 200

1st Stage Main Engine (4xRD0233) 1st stage boosterter 1st stage main engine (3 x RD-0233, 1 x RD-0234)

Figure 2-1 Rockot launch vehicle configuration.

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The booster unit, which provides the first plumes and gases, and ensures that the and second stages of Rockot, uses exist- correct temperature and humidity are main- ing SS-19 missiles and is accommodated tained during storage and operation. The within its existing transport and launch con- container is only used once. tainer. The third stage, which provides the orbital capability of the launcher, is newly 2.1.1 First Stage manufactured. This upper stage contains a modern, autonomous control and guidance The Rockot first stage has an external di- system which controls all three stages. The ameter of 2.5 metres and a length of 17.2 upper stage multiple engine ignition capa- metres. The main body of the stage con- bility allows implementation of various pay- tains N2O4 and UDMH tanks separated by load injection schemes. Figure 1-10 to Fig- a common bulkhead. Tank pressurisation ure 1-30 show the different components of is achieved by means of a hot gas system. the Rockot launch vehicle including the The engines are cardan-gimballed, closed- SS-19 booster stage contained within its cycle, turbopump-fed engines, three being transportation container, the payload fair- type RD-0233 and one type RD-0234. Fig- ing and the Breeze-KM upper stage. ure 1-3 shows a close-up view of the Rockot launch vehicle with the four en- Specifically, the Rockot launch vehicle gines ignited during lift-off. The first stage comprises: contains four solid fuel retro rockets for the first stage separation. • An existing SS-19 booster unit providing the 1st and 2nd stages The main stage characteristics are shown in Table 2-2 below. • An upper composite

The upper composite comprises: • Breeze-KM upper stage Main engines 3 x RD-0233, 1 x RD-0234 Propellant N2O4 / UDMH • Payload fairing 1870 kN Sea level thrust • Payload adapter or dispenser (each engine 470 kN) 2070 kN Vacuum thrust • Spacecraft (each engine 520 kN)

The launch takes place from the transport Sea level specific impulse 285 s and launch container erected above Vacuum specific impulse 310 s ground. The launcher rests physically on a Burn time 121 s ring at the bottom of the launch container. The umbilical between the launcher and Table 2-2 First stage main engine characteristics. the launch container is mechanically separated at lift-off. During lift-off, the launcher is guided by two rails within the launch container. The container protects the launch table environment from the engine

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2.1.2 Second Stage 2.1.3 Breeze-KM Upper Stage

The Rockot second stage has an external Figure 2-2 shows the Breeze-KM upper diameter of 2.5 metres and a length of 3.9 stage as part of the upper composite. In metres. It contains a closed-cycle, tur- addition to Breeze-KM, the upper compos- bopump-fed, fixed main engine designated ite consists of the payload fairing, the RD-0235 and vernier thrusters designated spacecraft adapter and the spacecraft. The spacecraft adapter is a mission specific RD-0236 for directional control. The four item, for details refer to Chapter 4. vernier thrusters have individual combus- tion chambers which are fed from one tur- The commercial version of Rockot uses the Breeze-KM as the standard upper stage bopump. Each thruster can be gimballed and is a close derivative of the original around one axis. The separation of the first Breeze-K flown during the first three and second stages is performed with the Rockot flights. It comprises three main vernier engines ignited just before the compartments which include the propulsion separation. The exhaust gases are di- compartment, the hermetically sealed verted by special hatches within the first equipment compartment and the interstage stage. After separation, the first stage is compartment. To allow larger satellites to decelerated by retro rockets before the be accommodated and to reduce dynamic second stage main engine is ignited. Like loads, structural changes to the Breeze-K the first stage it contains a common bulk- stage were introduced. The structure of the head and a hot gas pressurisation system. equipment bay of the original Breeze-K stage has been widened and flattened by a

redistribution of the control equipment. The equipment compartment can also dou- Main engine RD-0235 ble as a payload dispenser allowing multi- Propellant N2O4 / UDMH ple satellites to be easily accommodated. Vacuum thrust 240 kN Additionally, the compartment has been Vacuum specific impulse 320 s stiffened by the insertion of stiffening walls to give adequate structural rigidity. Fur- Burn time 183 s thermore, the Breeze-KM upper stage is no Table 2-3 Second stage main engine charac- longer attached to the launcher at its base teristics. but suspended within the extended interstage. The interstage is a load-bearing structure which provides the mechanical interface with the booster unit and accom- Vernier thrusters RD-0236 modates the Breeze-KM separation sys- (One turbopump and four thrusters) tem. Fuel N2O4 / UDMH Consequently, the fairing is attached di- Vacuum thrust in total 15.76 kN rectly to the equipment compartment. A large variety of different payload configura- Vacuum specific impulse 293 s tions can be accommodated, ranging from Burn time 200 s single to multiple satellite launches, posi- Table 2-4 Second stage vernier thrusters tioned either on a single level or on two or characteristics. more levels using a customised dispenser.

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Payload fairing static envelope

Payload adapter

Payload fairing Payload separation plane Equipment compartment / dispenser

Interstage

Breeze-KM Upper composite upper stage installation plane

Figure 2-2 Upper composite with payload.

The Breeze-KM equipment compartment system has three independent chan- contains: nels with majority voting and is totally autonomous with respect to ground • A telemetry system including transmit- control. ters and antennas. Breeze-KM also contains tape recorders for store-and- • A tracking system with receiver/ trans- forward telemetry capability. mitter and antennas

• A guidance, navigation and control sys- The Breeze-KM can be equipped with two tem for all flight phases and manoeu- or three batteries which can supply both vres before and after spacecraft sepa- Breeze-KM and payload systems, see sec- ration. It contains an inertial guidance tion 4.3.4.2 for further details. The propul- system based on a 3-axis gyro platform sion compartment consists of fuel com- with an on-board computer. The control partment and engines including associated equipment.

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The Breeze-KM propulsion compartment Type Bipropellant pressure fed contains propellant tanks and the propul- Vacuum thrust (each) 400 N sion systems. The low pressure fuel tank Total available impulse 141120 Ns (UDMH) and the oxidiser tank (N2O4) are Minimum impulse bit 40 Ns separated by a common bulkhead. The Operation mode Pulse or steady state Polyus, Kvant.2, Krystall, Previous flight heritage lower oxidiser tank surrounds the 20 kN Spectr, Priroda, FGB. main engine. Each tank contains equip- Table 2-6 Breeze-KM vernier engine charac- ment such as baffles, feed pipes and ul- teristics. lage control devices to facilitate main engine restarts in weightlessness. Type Bipropellant pressure fed The Breeze-KM propulsion system in- Vacuum thrust (each) 13 N cludes the main engine, the attitude and Total available impulse - orbit control system (AOCS) and vernier Minimum impulse bit 0.068 Ns thrusters together with propellant feed lines Operation mode Pulse or steady state and spherical nitrogen gas tanks. The 12 x Previous flight heritage Polyus, Kvant.2, Krystall, Spectr, Priroda, FGB. 13 N attitude control engines control the Table 2-7 Breeze-KM AOCS engine character- pitch, roll and yaw degree of freedom of istics. the Breeze-KM vehicle. The 4 x 400 N vernier thrusters which are located at the base of the Breeze-KM are used for pro- Component Mass, kg pellant settling and orbital manoeuvres. Dry mass 1320 Oxidiser N2O4 max 3310 The 20 kN main engine provides the major Propellant UDMH max 1665 impulse required to achieve the final orbit. Table 2-8 Estimated Breeze-KM mass break- The characteristics and extensive flight down. heritage of all these engines are shown in Table 2-5 to Table 2-7. Table 2-8 summa- rizes the mass breakdown of Breeze-KM. 2.1.4 Fairing

The payload fairing has been specially designed for the commercial version of Closed cycle turbo Type Rockot and is based on proven technology pump fed from other KSRC programmes. Vacuum thrust 20 kN Vacuum specific impulse 325.5 s The fairing is mounted on top of the Number of ignitions up to 8 equipment bay of the upper stage. The Total available impulse 2 x 107 Ns fairing separation and jettison are Minimum impulse bit 25000 Ns performed by releasing mechanical locks Maximum burn time 1000 s Minimum burn time 1 s holding the two half-shells together along Off time 15 s to 5 hours the vertical split line via a pyrodriver lo- Phobos-1, Phobos-2 cated in the nose of the fairing. This py- Previous flight heritage and Mars 91 space rodriver has redundant firing circuits. Im- vehicles. mediately following the release of the Table 2-5 Breeze-KM upper stage main en- locks, several pyrobolts on the fairing’s gine characteristics.

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horizontal split line are fired to allow the • An extension for the upper stage and half-shells to be driven sideward by spring payload fairing pushers. The half-shells rotate around • Internal guiding rails hinges located at their base and are sub- sequently jettisoned. • Systems for fuelling, pressurisation thermal control and electrical support The design concept is based on the current commercial Proton fairing design. The fair- 2.2 Rockot Qualification and ing is fabricated from a three layer carbon Flight History fibre composite with an aluminium honey- comb core. KSRC has been using these The Rockot launch system has a long flight materials for payload fairings since 1985. heritage with an excellent record. To main- They are especially suitable for absorbing tain this impressive track record, which acoustic noise. includes an unbroken run of over 80 launches of the Rockot booster stage (SS- The fairing separation system has an ex- 19) without launch failure since 1983. cellent design heritage. Its mechanisms EUROCKOT has purposely retained as have been extensively ground- tested and much of this heritage as possible in its successfully used in numerous flights of commercial version of the vehicle. different KSRC programmes. Rockot's first three launches took place The payload fairing dynamic envelope is with the Rockot/Breeze-K configuration given in chapter 4. and were launched with a small fairing from a silo in the Baikonur Cosmodrome. 2.1.5 Transport and Launch Container Launches one and two were performed on 20th November 1990 and 20th December A special feature of the Rockot launch ve- 1991, respectively. Geophysical experi- hicle is the use of a transport and launch ments were performed during these flights. container (TLC), providing the following During these launches, after first and sec- functions: ond stage burn-out, separation of the up- • Storage of booster unit under climati- per stage Breeze-KM from the second cally controlled conditions stage booster was successfully performed and a sub-orbital controlled and stabilised • Booster unit transportation flight of the upper stage, which carried the • Launch vehicle erection on pad scientific equipment, was undertaken to a maximum altitude of 900 km and inclination • Launch vehicle pre-launch preparation and environmental protection of 65°.

• Launch vehicle physical guidance dur- Multiple restarts of the upper stage main ing lift-off engine were performed during every flight. The first launches permitted testing of the It consists of: efficiency of all the launch vehicle’s equip- • A cylindrical container ment and systems, estimation of the upper stage dynamic performance in weightless

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conditions during the propulsion unit multi- All modifications underwent a thorough ple restarts, and acquisition of the data on ground qualification program prior to the levels of shock, vibrational and acoustic first launches. loads. All following launches were performed The third launch of Rockot was success- using the commercialised Rockot/Breeze- fully performed on 26th December 1994. KM version. They were prepared and con- As a result of this launch, the Radio- ducted from EUROCKOT's dedicated ROSTO radio-amateur satellite having a launch pad and facilities in Plesetsk Cos- mass of about 100 kg was injected into a modrome. The first launch to be performed circular orbit of 1900 km at an inclination of under EUROCKOT management was the 65°. Multiple restarts of the upper stage Commercial Demonstration Flight (CDF) main engine were also performed during which injected two satellite simulators this flight. SIMSAT-1 and SIMSAT-2 extremely accurately into their intended orbit. The CDF Since the Rockot/Breeze-K configuration launch enabled the following objectives to launched from Baikonur could not ade- be demonstrated: quately serve the high and polar inclination market identified by EUROCKOT and, fur- • Readiness of Rockot installations in thermore, did not allow large LEO payloads Plesetsk for commercial operations to be accommodated within the existing • Provision of flight verification of the envelope, EUROCKOT modified the Rockot/Breeze-KM configuration Rockot/Breeze-K launch vehicle for com- • Injection of two satellite simulators mercial operations and opened up a new SIMSAT-1 and SIMSAT-2 into a 547 launch base at Plesetsk Cosmodrome in km circular orbit at 86.4° inclination Northern Russia. To retain the heritage of • Testing and verification of technical Rockot/Breeze-K and SS-19 missile facilities, the launch pad, fuelling sys- launches from the silo-based TLC, an iden- tems, operations, electrical ground tical system of launching from a container support equipment, and data meas- is used for Rockot/Breeze-KM version urement, recording and processing sys- launched from above the ground at Ple- tems setsk Cosmodrome. Similarly, no major • Measurement and evaluation of the systems such as the vehicle avionics and payload environment during flight and control system or propulsion have been confirmation of User's Guide data modified for the commercial Rockot/ • Demonstration of the Rockot launch Breeze-KM launcher, only structural vehicle system's inherent reliability changes to the upper composite have been A full list of the payloads launched by made (section 2.1.3). Rockot/Breeze-KM is shown in Table 2-9.

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No. Payload Date Comments

Commercial Demonstration Flight Rockot/Breeze-KM from Plesetsk, 1 16.05.00 [EUROCKOT, Germany] operated by EUROCKOT, success

GRACE 1, GRACE 2 Rockot/Breeze-KM, operated by 2 17.04.02 [German Aerospace Centre DLR and National EUROCKOT, success Aeronautics & Space Administration, USA]

Iridium SV97, Iridium SV98 Rockot/Breeze-KM, operated by 3 20.06.02 [Iridium Satellite LLC, USA] EUROCKOT, success

Multiple Orbit Mission (MOST, MIMOSA, 6 Nanosatellites Rockot/Breeze-KM, operated by 4 [Canadian Space Agency, Czech Astronomical 30.06.03 Institute and different research institutes and EUROCKOT, success universities, Czech Republic, Canada, Japan, Denmark and USA] SERVIS-1 Rockot/Breeze-KM, operated by 5 30.10.03 [Institute for Unmanned Space Experiment Free EUROCKOT, success Flyer, Japan]

CryoSat Rockot/Breeze-KM, operated by 6 08.10.05 [European Space Agency] EUROCKOT, failure

KOMPSAT-2 Rockot/Breeze-KM, operated by 7 28.07.06 [Korean Aerospace Research Institute KARI, EUROCKOT, success South Korea]

GOCE Rockot/Breeze-KM, operated by 8 17.03.09 [European Space Agency] EUROCKOT, success

SMOS, PROBA-2 Rockot/Breeze-KM, operated by 9 02.11.09 [European Space Agency] EUROCKOT, success SERVIS-2 Rockot/Breeze-KM, operated by 10 02.06.10 [Institute for Unmanned Space Experiment Free EUROCKOT, success Flyer, Japan] Additionally, three successful Rockot/Breeze-K launches were performed in the early 1990s. In parallel to EUROCKOT activities starting in 2005 the actual Rockot/Breeze-KM version has been used for five Russian federal launches, of which four were successful. For further information on details of these launches and the achieved injection accuracies please contact EUROCKOT directly.

Table 2-9 EUROCKOT launch record with Rockot/Breeze-KM.

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2.3 Revalidation of SS-19s used by EUROCKOT

The SS-19 booster units used by

EUROCKOT for the Rockot launch vehicle are existing ICBM assets which have been assigned to EUROCKOT by the Russian government. SS-19s received by KSRC principally undergo a revalidation programme prior to being used for the Rockot launch vehicle. The revalidation procedure is beyond the scope of this User's Guide.

However, the procedure includes the following steps:

• After draining of the fuel, the SS-19s are removed from their silos for storage

• The SS-19s are stored under climatically controlled conditions in a defu- elled state within their transport containers until the beginning of launch preparations. The atmosphere within the containers is climatically controlled

at all times using dry nitrogen gas.

• Constant quality control checks of stored batches of SS-19s via a regular test programme which involves subject- ing parts of the batches to flight tests, engine hot firing tests and destructive physical analyses including metallurgi- cal tests as well as functional tests on the stored boosters.

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Chapter 3 General Performance Capabilities

Table of Contents

3 General Performance Capabilities...... 3-1 3.1 Introduction...... 3-1 3.2 Launch Azimuths and Orbit Inclinations from Plesetsk ...... 3-1 3.3 Low Earth Orbits...... 3-2 3.3.1 Payload Performance for Circular Orbits ...... 3-2 3.3.2 Performance for Elliptical Orbits ...... 3-3 3.3.3 Sun-Synchronous Orbits...... 3-3 3.4 Mission Profile Description ...... 3-8 3.5 Baikonur Performance...... 3-13 3.6 Spacecraft Injection and Separation ...... 3-14 3.6.1 Injection Accuracy...... 3-14 3.6.2 Spacecraft Separation ...... 3-14 3.6.2.1 Spin Stabilised Separation...... 3-14 3.6.2.2 Three-Axis Stabilised Separation...... 3-15 3.6.2.3 Typical Multiple Satellite Deployment Scenarios ...... 3-16

List of Figures

Figure 3-1 Rockot/Breeze-KM payload performance for circular orbits...... 3-4 Figure 3-2 Rockot/Breeze-KM payload performance for elliptical orbits at 63.2° inclination...... 3-5 Figure 3-3 Rockot/Breeze-KM payload performance for elliptical orbits at 72.0° inclination...... 3-6 Figure 3-4 Rockot/Breeze-KM payload performance for elliptical orbits at 82.5° inclination...... 3-7 Figure 3-5 Cycling orientation of Breeze-KM relative to the sun during coast flight maintained for one hour (a), and for half an hour (b)...... 3-9 Figure 3-6 Typical Rockot/Breeze-KM ascent (650 km circular, 86.583° inclination). (a) Trajectory. (b) Acceleration timeline...... 3-10 Figure 3-7 Typical Rockot/Breeze-KM ascent (1000 km circular, 99.52° inclination). (a) Trajectory. (b) Acceleration timeline...... 3-11 Figure 3-8 Typical Rockot/Breeze-KM ascent (320 km perigee, 820 km apogee at 96.8° inclination and SSO 820 km, at 98.7° inclination). (a) Trajectory. (b) Acceleration timeline...... 3-12 Figure 3-9 Sun-synchronous orbit injection scheme...... 3-13 Figure 3-10 Example of sequential deployment of three spacecraft...... 3-16 Figure 3-11 Example of simultaneous deployment of six spacecraft...... 3-17

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List of Tables

Table 3-1 Allowable launch azimuths that can be served from Plesetsk...... 3-1 Table 3-2 Orbital injection accuracy...... 3-15

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3 General 3.2 Launch Azimuths and Orbit Inclinations from Plesetsk Performance The Plesetsk Cosmodrome is located Capabilities about 200 km south of the port city of Archangel in Northern Russia at geo- This chapter describes the performance of graphical coordinates 62.8°N and 40.3°E. the Rockot/Breeze-KM launch vehicle into The location of populated areas drives the circular and elliptical low earth orbits from allowable launch azimuths and drop zones its launch site in the Plesetsk Cos- available from this launch site. Launch modrome, Northern Russia, as well as its azimuths and resulting orbital inclinations potential launch base in Baikonur. Back- achievable from Plesetsk are listed in ground information and the assumptions Table 3-1. made for the performance curves are pre- Launch Azimuth Corresponding Orbital Inclination sented. 76.2° 63.2° 41.6° 72.0° 3.1 Introduction 13.7° 82.5° The launch vehicle payload performance 15.2° to 4.8° 82.0° to 86.4° is driven by many variables and includes 4.8° 86.4° amongst others the specific launch vehicle 341.5° SSO and other retrograde orbits characteristics, launch pad location, al- Table 3-1 Allowable launch azimuths that lowable launch azimuths, drop zones, and can be served from Plesetsk. the availability of ground measuring stations for telemetry information reception. Rockot/Breeze-KM, equipped with its The Rockot launch site in the Cos- modern inertial based control system is modrome Plesetsk, historically the most able to perform dog-leg manoeuvres dur- active launch site in the world with over ing the second stage operation so that 1700 launches, is well situated for polar inclinations beyond these allowable and high inclination launches due to its launch azimuths can be reached. The northerly latitude. Rockot/Breeze-KM dog-leg manoeuvres may result in a de- launched from Plesetsk Cosmodrome and crease of payload mass. equipped with its modern restartable upper stage Breeze-KM can serve a wide In coordination with the Customer and range of both circular and elliptical orbits their demands the Breeze-KM upper stage in the range from 200 km up to over 2000 enables high flexibility in the selection of km and a range of inclinations from 50° to the ascent profile provided by its attitude- SSO by direct injection or via orbital plane and orbit correction system, precise change. guidance, navigation and control electron- ics including a three-axis gyro system and long life batteries. This enables a Cus- tomer adapted ascent profile and payload

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deployment scheme under consideration parking orbit with the first burn of the of radiovisibility by Russian ground track- main engine and one or several ad- ing stations, earth shadow phases, sepa- justment burns to form the target orbit. ration time and other constraints. Note: If the required altitude of the orbit To achieve inclinations other than those does not exceed 400 km, both injection indicated in Table 3-1, Breeze-KM also schemes can be used, and if the orbit alti- provides the possibility to change inclina- tude is higher than 400 km, the second tion up to ±17° by a main engine ignition injection scheme is generally used. in the vicinity of the equatorial node of the All payload performances are calculated transfer orbit. In such cases the possible for the standard Rockot/Breeze-KM decrease of the payload mass has to be configuration including the payload fairing determined for each specific mission pro- as described in chapter 2. The requisite file. The minimum possible orbital inclina- payload adapter fitting/ dispenser masses tion for the launches from Plesetsk cos- plus the separation system must be sub- modrome without dog-leg manoeuvres tracted from these figures. and/or main engine ignition in the vicinity of the nodes is 62.8°. Usually, the payload fairing is not jettisoned until the free molecular heat-flow The propellant consumed by Breeze-KM has dropped below 1135 W/m2. during possible payload collision and contamination avoidance manoeuvres is mi- The performance values are confirmed by nor and will not affect the payload per- the data of former Rockot/Breeze-KM formance. On the other hand, fuel con- commercial launches as well as the over sumption for possible Breeze-KM deorbi- 150 SS-19 missile flights. tation must be subtracted from the performance capacity. It should be noted that the performances given in this user guide are generally 3.3 Low Earth Orbits based on conservative assumptions. Fur- thermore, due to mass saving measures The payload performance of the such as incremental improvements to the Rockot/Breeze-KM vehicle has been cal- upper stage, an increase in payload per- culated for both circular and elliptical or- formance can be expected. In specific bits from the Plesetsk launch site. To at- cases, where such additional performance tain the maximum payload capacity for a is necessary, the Customer is invited to dedicated mission, two injection schemes contact EUROCKOT directly for a dedi- are generally used: cated mission analysis.

• The target orbit is achieved via a sin- 3.3.1 Payload Performance for Circular gle burn of the Breeze-KM upper Orbits stage main engine. Figure 3-1 illustrates the performance ca- • The Breeze-KM upper stage with a pabilities associated with the correspond- payload is injected into an elliptic ing circular orbits that can be served from

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the launch site in Plesetsk using the al- 3.3.3 Sun-Synchronous Orbits lowable launch azimuths indicated in sec- Sun-synchronous orbits (SSO) can be tion 3.2. It should be noted that direct in- served from the Plesetsk launch site via jection into inclinations that lie between use of the 341.5° launch azimuth corridor. i = 82.0° and 86.4° are possible but Different ascent trajectory options are subject to a dedicated internal Russian available depending on the requirements approval process for the overflight permis- of the dedicated mission. sion. Inclinations not shown in the performance graphs can also be served by The launch vehicle is initially launched Rockot/Breeze-KM, but only via a dog-leg with an azimuth of 341.5° from Plesetsk. manoeuvre during the 2nd stage burn or a Yaw manoeuvres during the second stage plane change manoeuvre performed by burn allow the second stage drop zone to the upper stage. In these cases, perform- be precisely positioned outside of any for- ances should be calculated on a case by eign country's territorial waters. case basis by EUROCKOT, a linear inter- polation between the curves is not possi- The upper composite comprised of ble. Some loss of performance can be Breeze-KM and the payload is injected expected due to the necessity to perform into a 96.7° or 99.5° inclined parking orbit. dog-leg or plane change manoeuvres. Finally, the target orbit inclination is reached via a plane change manoeuvre 3.3.2 Performance for Elliptical Orbits carried out by a Breeze-KM main engine ignition near the equator crossing. The Rockot/Breeze-KM performance ca- The payload performance for SSO is de- pabilities for elliptical orbits with inclina- picted in Figure 3-1. It corresponds to the tions of 63.2°, 72.0° and 82.5° are pre- payload capacity into the required orbit sented in Figure 3-2 to Figure 3-4. Any with the SSO typical combination of target argument of perigee can be achieved ac- altitude and inclination. cording to the Customer's requirements.

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2200

2000

1800 i = 63.2° i = 72.0° i = 82.5° 1600 i = 86.4°

1400

1200 Payload mass, kg Payload mass, SSO Parking orbit i = 96.7° 1000 Parking orbit i = 99.5°

800

600

400 0 500 1000 1500 2000

Circular orbit altitude, km

Figure 3-1 Rockot/Breeze-KM payload performance for circular orbits.

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2000 Perigee altitude, km 200 500 700 1000 1500 1500 2000

1000 Payload mass, kg Payload mass,

500

0 100 1000 10000 100000

Apogee altitude, km

Figure 3-2 Rockot/Breeze-KM payload performance for elliptical orbits at 63.2° inclination.

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2000

Perigee altitude, km 200 500

1500 700 1000 1500 2000

1000 Payload mass, kg Payload mass,

500

0 100 1000 10000 100000

Apogee altitude, km

Figure 3-3 Rockot/Breeze-KM payload performance for elliptical orbits at 72.0° inclination.

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2000

Perigee altitude, km 200 500 700 1000 1500 1500 2000

1000 Payload mass, kg Payload mass,

500

0 100 1000 10000 100000

Apogee altitude, km

Figure 3-4 Rockot/Breeze-KM payload performance for elliptical orbits at 82.5° inclination.

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the Customer’s specific requirements be- 3.4 Mission Profile Description tween the main engine burns. For instance, This section describes typical circular low- for thermal sensitive payloads, a step-wise earth mission profiles and presents exam- rotation about any axis can be performed. ples of trajectories. In the absence of any specific requirements, Breeze-KM performs manoeuvres The selected flight trajectories take into to meet its own thermal requirements. For account the dedicated impact sites permit- 1 hour the +XUS axis is oriented towards ted for Rockot elements. the sun. If coasting continues for more than

one hour, the –XUS axis is oriented towards The launch sequence begins with Stage 1 the sun for the next half an hour. During ignition. The first stage propels the vehicle +XUS orientation, the angle between the to approximately 60 km height and impacts +XUS axis and the direction of sun light some 990 to 1100 km down range. The should be α ≤ 100°. During –XUS orienta- ignition of the Stage 2 vernier engines oc- tion, the angle between the –XUS axis and curs shortly before Stage 1 burn-out. the sun direction should be α ≤ 50° (Figure After shut down of stage 1 engine, stage 1 3-5). is separated using its solid retro rockets. For a chosen orientation, an accuracy of Once the free molecular heat-flow has 10° about each of the three axes of stabili- fallen below 1135 W/m2, usually, the pay- sation can be provided during the coast load fairing is jettisoned. phase. The second stage's propelled flight phase Typical trajectories for Rockot missions is completed by the successive shut down with different numbers of upper stage igni- of the main engine and vernier thrusters. tions are shown in Figure 3-6 to Figure 3-8. The following stage separation is assisted These figures show the main events of the by use of the second stage's retro rockets. mission: main engine burns and cut-offs, The Breeze-KM upper stage manoeuvres the Spacecraft separation and trajectory begin immediately after stage 2 separation characteristics, such as: and are performed by the upper stage • flight time t counted from launch, s main engine, which can be ignited several times, if required. An initial burn is per- • relative velocity v, m/s formed in the boost mode directly following stage 2 separation. Further ignitions of the • relative flight path angle θ, deg main engine are performed in accordance • dynamic pressure q, kg/m2 with the specific flight programme. • altitude h, km During the coast phase between the main engine burns Breeze-KM generally follows Figure 3-9 shows a typical injection a sun-oriented flight programme to meet scheme for sun-synchronous orbits.

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-50°≤α≤ -100 ≤ α ≤ 100 -X ++X XUS - XUS α ≤ 50° α ≤ 100° α α U Breeze-KM PayloadPayload B Directionction from from Directionction from from GCoG to Su ton sun GCoG to Sunto sun Upper Stage Breeze-KMBreeze-KM Centreter ofof Gravitygravity Payload Centrenter ofof Gravitygravity G) (CoG) OG) (CoG)

per StagBreeze-Ke M er Breeze-KStage M longitudinalgitudinal axi axiss longitudinalgitudinal axisaxis

(a) (b)

Figure 3-5 Cycling orientation of Breeze-KM relative to the sun during coast flight maintained for one hour (a), and for half an hour (b).

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Spacecraft End of deployment Breeze-KM t = 5861 s second burn Orbit parameters t = 4850 s h = 650.0 km v = 7622 m/s p ha = 650.0 km Stage 2 θ = 0.02° Fairing separation i = 86.583° separation, t = 173 s Final orbit Breeze-KM parameters q = 0 Pa first burn v = 3599 m/s hp = 499.2 km t = 305 s θ = 16° ha = 506.1 km v = 5634 m/s h = 125.6 km i = 88.999 θ = 7.3° h = 235.7 km Breeze-KM second burn Breeze-KM Stage 1 withdrawal separation t = 4360 s v = 7367 m/s t = 6915 s t = 122 s θ = -0.03° q = 586.44 Pa h = 665.6 km Orbit parameters v = 3163 m/s End of Breeze-KM hp = 140.7 km θ = 23.3° first burn ha = 660.8 km h = 68.3 km t = 882 s i = 86.582° v = 7832 m/s θ = -1.5° h = 250.7 km Maximum Transfer orbit dynamic pressure parameters

t = 50 s hp = 160 km q = 61.29 kPa ha = 635.3 km v = 567 m/s i = 86.577° Launch θ = 47.9° h = 10.7 km t = 0 s L = 1076 km L = 1391 km L = 3662 km

time

(a)

Stage 1 and 2 Operation Breeze Operation Transfer Orbit Stage 1 Stage 2 Separation Separation Breeze Spacecraft Breeze Fairing Deployment Venting/ Jettisoning Breeze Second Withdrawal Depressurisation First Burn Burn of Breeze Tank

Launch 0 122 173 305 882±43 4360 4399±3 5861 6915 t (s Retro Rocket Retro Rocket time, s of Stage 1 of Stage 2

(b)

Figure 3-6 Typical Rockot/Breeze-KM ascent (650 km circular, 86.583° inclination). (a) Trajectory. (b) Acceleration timeline.

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Spacecraft End of deployment Breeze-KM t = 5768 s second burn t = 4505 s Orbit parameters v = 7455 m/s hp = 1000.0 km θ = -0.006° ha = 1000.0 km i = 99.52° Final orbit parameters Breeze-KM withdrawal hp = 1001.1 km ha = 1009.5 km t = 6368 s Stage 2 i = 99.516 Fairing separation separation, Orbit parameters t = 164 s Breeze-KM hp = 776.5 km q = 0 Pa first burn ha = 1001.1 km v = 3480 m/s t = 305 s i = 99.594° θ = 18.2° v = 5543 m/s h = 119.9 km θ = 9.7° h = 257.2 km Breeze-KM second burn Stage 1 separation t = 4456 s v = 7226 m/s t = 122 s θ = 0.07° q = 479.55 Pa h = 1005.9 v = 3153 m/s End of Breeze-KM θ = 24.2° first burn h = 69.7 km t = 922 s v = 7984 m/s θ = -2.4° h = 291.3 km Transfer orbit Maximum parameters dynamic pressure h = 160 km t = 50 s p ha = 976.3 km q = 60.63 kPa i = 99.46° v = 566 m/s θ = 48.8° Launch h = 10.7 km t = 0 s L = 1082 km L = 1329 km L = 3857 km

time

(a)

Stage 1 and 2 Operation Breeze Operation Transfer Final Orbit Orbit Stage 1 Stage 2 Separation Separation Breeze Spacecraft Fairing Deployment Breeze Venting/ Jettisoning Breeze Second Depressurisation First Burn Burn Withdrawal of Breeze Tank

Launch 0 122 164 305 922±45 4456 4505±3 5768 6368 t ( Retro Rocket Retro Rocket time, s of Stage 1 of Stage 2

(b)

Figure 3-7 Typical Rockot/Breeze-KM ascent (1000 km circular, 99.52° inclination). (a) Trajectory. (b) Acceleration timeline.

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End of Breeze-KM third burn Spacecraft t = 5213 s deployment v = 7537 m/s θ = -0.006° t = 5481 s Final orbit Orbit parameters parameters hp = 820.0 km h = 820.0 km hp = 820.1 km a h = 828.9 km i = 98.7° Spacecraft a i = 98.696° deployment Fairing separation t = 2826 s Stage 2 t = 171 s separation, Breeze-KM Orbit q = 0 Pa Breeze-KM Breeze-KM second burn parameters v = 3638 m/s first burn third burn t = 2636 s hp = 320.0 km θ = 14.3° t = 5156 s t = 305 s v = 7785 m/s ha = 820.0 km h = 120.2 km v = 7393 m/s v = 5892 m/s θ = 1.41° i = 96.80° θ = 0.86° θ = -3.8° h = 326.9 km h = 200.4 km h = 820.9 km Stage 1 separation t = 122 s End of Breeze-KM Breeze-KM q = 653.12 Pa second burn withdrawal v = 3212 m/s End of Breeze-KM t = 2687 s t = 6061 s θ = 22.6° first burn v = 7917 m/s h = 67.7 km t = 800 s θ = -0.0004° Orbit parameters v = 7920 m/s h = 208.8 km hp = 673.7 km θ = -0.08° ha = 821.1 km h = 208.8 km Transfer orbit i = 98.703° parameters Maximum Transfer orbit hp = 320.6 km dynamic pressure parameters ha = 826.9 km t = 50 s hp = 160 km i = 96.80° q = 62.65 kPa ha = 470.0 km v = 574 m/s Launch i = 96.864° θ = 47.1° t = 0 s h = 10.7 km L = 1082 km L = 1325 km L = 3201 km

time

(a)

Stage 1 and 2 Operation Breeze Operation Transfer Final Orbit Final Orbit Orbit Stage 1 Stage 2 Separation Separation Breeze Spacecraft Breeze Spacecraft Fairing Deployment Third Deployment Breeze Venting / Jettisoning Breeze Second Depressurisation First Burn Burn Burn Withdrawal of Breeze Tank

Launch 0 122 171 305 799±36 2636 2688±32826 5156 5213±3 5481 6061 t (se Retro Rocket Retro Rocket of Stage 1 of Stage 2 time, s

(b)

Figure 3-8 Typical Rockot/Breeze-KM ascent (320 km perigee, 820 km apogee at 96.8° inclination and SSO 820 km, at 98.7° inclination). (a) Trajectory. (b) Acceleration timeline.

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Required orbit

Stage 2 separation End of ascent phase Fairing separation

Coast phase Stage 1 separation

Descending node. Breeze-KM engine ignition

Ascending node. Breeze-KMBREEZE engine ignition

Transfer orbit Equatorial plane

Figure 3-9 Sun-synchronous orbit injection scheme.

missions. Baikonur is particularly suited for 3.5 Baikonur Performance serving inclinations in the 50° range, which Although Rockot launches have occurred cannot be efficiently reached from Plesetsk in the past from Baikonur cosmodrome, it due to its northerly latitude. The achievable is currently no longer operational for mass performances from Baikonur for in- Rockot launches. An activation of the site clinations from about 50° to about 65° are for Rockot requires extensive upgrades subject to detailed analyses. However, and modifications to existing facilities. A they can be assumed to be slightly above decision for site activation will be made on the mass performance for launches from a case-by-case basis, should the need Plesetsk. For more detailed information arise. The information contained in this regarding Baikonur launch performance, section is for customers interested in the EUROCKOT should be contacted, directly. potential use of this launch site for their

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3.6 Spacecraft Injection and Spacecraft can either be spun-up along the Separation Breeze-KM longitudinal axis or released from the upper stage in a three-axis stabi- Rockot, equipped with the Breeze-KM up- lised mode. per stage allows a large variety of options with regard to spacecraft orbital injection 3.6.2.1 Spin Stabilised Separation and separation. The following sections provide information about the orbital injec- Spin stabilisation is performed around the tion conditions and the separation possibili- upper stage longitudinal axis with rates of ties for payloads. up to 10°/min. Higher spin rates may be considered upon Customer's request. 3.6.1 Injection Accuracy Spin parameters are to be agreed Table 3-2 provides 3-sigma orbital injection individually for each specific payload taking errors depending on the average altitude of into account: the target orbit ensured by accuracy prop- • Payload mass distribution (MoI), centre erties of the Rockot/Breeze-KM launch of mass (CoM) position and spacecraft vehicle control system. dynamic properties In particular cases the injection accuracies • Customer requirements for the spin given in Table 3-2 can be improved after regime such as: the specific trajectory analysis. − attitude orientation and its accuracy during upper stage spin manoeuvre 3.6.2 Spacecraft Separation − orientation accuracy of the payload af- The spacecraft separation from Breeze-KM ter its separation can take place in a number of different − other payload requirements for the ways and is driven primarily by the Breeze-KM upper stage • Necessity to continue flight control of • characteristics of the separation sys- the upper stage after payload separa- tem, e.g. stiffness of spring pushers, tion • type of release mechanism, Controlled deorbiting of the upper stage • the direction of the separation impulse after separation can be provided, if re- of the payload, quired. On the completion of the mission, Breeze-KM vents all its tanks to put the payload mass, moments of inertia, • stage in a safe mode. • the Breeze-KM burn-out mass

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Orbital parameters error type 3-Sigma error

Average orbital altitude ±1.5%

Inclination ±0.06° Eccentricity (for circular orbits) ≤0.0025

Right ascension of ascending node ±0.05° Argument of perigee (for elliptical orbits) depends on eccentricity

Table 3-2 Orbital injection accuracy.

3.6.2.2 Three-Axis Stabilised Separation The schemes are selected by EUROCKOT and agreed with the Customer. In general, any required payload attitude can be provided. Following orbit insertion, As a possible way to reduce potential dis- the Breeze-KM avionics subsystem can turbances acting on the spacecraft during execute a series of pre-programmed com- separation, the following measures shall be mands to provide the desired initial pay- applied: load attitude prior to its separation. • Selection of pusher with optimal char- This capability can also be used to reorient acteristics including their energy Breeze-KM for the deployment of multiple payloads with independent attitude re- • Adjustment of pusher position for com- quirements. pensation of lateral offset of the CoM

The 3-sigma attitude error about each Besides, electrical connectors can be se- spacecraft geometrical axis will not exceed lected taking into account their separation 1.5° - 3° depending on the spacecraft prop- forces characteristics. erties. Analysis shows that even for light space- The maximum angular velocities of the craft having a mass of not more than 500 combined Breeze-KM / spacecraft prior to kg and moments of inertia of up to 50 2 the payload deployment are: kg·m the total angular velocities ωy and ωz will not exceed 2.5°/s and the longitudinal

ωx = ± 1°/s component of ωx will not exceed 1.5°/s af- ωy = ± 0.5°/s ter separation, if the above methods are ωz = ± 0.5°/s combined. For spacecraft of larger size and higher inertia the disturbance values The spacecraft separation scheme system are smaller. design including the number of pushers, their allocation and energy is developed in Please note that these values shall be accordance with the requirements for en- considered as conservative ones. The ac- suring the spacecraft normal operations. tual parameters can be smaller depending on the properties of the specific spacecraft.

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To fulfil the Customer's requirements for Depicted in the figures below are two typi- the spacecraft separation, EUROCKOT cal payload deployment schemes. Figure selects the best suited of above methods 3-10 shows the sequential separation of to provide an optimal solution. three spacecraft with Δv added to each spacecraft to aid in-orbit plane phasing. 3.6.2.3 Typical Multiple Satellite Deployment Figure 3-11 shows a sequence in which six Scenarios spacecraft are released simultaneously.

Breeze-KM is able to perform a wide vari- The separation scenario of the spacecraft ety of complex pre-programmed manoeu- is laid out in accordance with number, ar- vres using a combination of its main, rangement and energy of the pushers and vernier and attitude control engines, that their requirements for normal operation of allow to implement injection of several pay- the satellite. The separation scenario is loads into specified target orbits. selected in cooperation with the Customer.

SC separation velocity 3 m/s SC separation velocity 3 m/s 26.5 26.5

t = 40 s t = 40 s Turn to Turn to t = 60 s payload t = 60 s t = 45 s t = 10 s payload Release SC2 separation Release SC1 and Realign Breeze-KM Impulse separation and stabilise direction stabilise Breeze-KM with velocity vector Δv = 5 m/s direction Breeze-KM

SC separation velocity 3 m/s

26.5

Target orbit t = 40 s plane Turn to t = 100 s t = 45 s t = 10 s payload t = 60 s Reorient Breeze-KM Realign Breeze-KM Impulse separation Release SC3 and Breeze-KM for deorbit with velocity vector Δv = 5 m/s direction stabilise Breeze-KM deorbiting manoeuvre

Figure 3-10 Example of sequential deployment of three spacecraft.

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TargetTarget Orbit orbit Planeplane

Upper composite Rotation around Rotation along the upper aligned with velocity the pitch axis stage longitudinal axis vector followed by simultaneous release of satellites into the target orbit plane

Figure 3-11 Example of simultaneous deployment of six spacecraft.

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Chapter 4 Spacecraft Interfaces

Table of Contents

4 Spacecraft Interfaces...... 4-1 4.1 Mechanical Interfaces ...... 4-1 4.1.1 Payload Accommodation ...... 4-1 4.1.2 Usable Volume for Payload ...... 4-3 4.1.3 Spacecraft Accessibility ...... 4-3 4.1.4 Separation Systems...... 4-3 4.1.4.1 Clamp Band Separation Systems...... 4-5 4.1.4.2 Mechanical Lock Systems ...... 4-9 4.1.4.3 Minisatellite Separation System ...... 4-10 4.1.5 Payload Adapters ...... 4-12 4.1.5.1 Clamp Band Separation System Adapters ...... 4-12 4.1.5.2 Mechanical Lock System Payload Adapters ...... 4-17 4.2 Electrical Interfaces ...... 4-20 4.2.1 On-board Interfaces...... 4-20 4.2.1.1 Umbilical Connectors...... 4-20 4.2.1.2 Separation Verification ...... 4-20 4.2.1.3 Interface Electrical Constraints...... 4-20 4.2.1.4 Umbilical Harness Configuration and Specifications ...... 4-22 4.2.1.5 Matchmate / Electrical Checkout ...... 4-24 4.2.1.6 Spacecraft Electrical Interface Input Data Requirements...... 4-24 4.2.2 Ground Electrical Interface ...... 4-24 4.2.3 Payload Grounding and Bonding...... 4-26 4.2.4 Payload Auxiliary Power Supply ...... 4-28 4.2.4.1 Ground Auxiliary Power Supply...... 4-28 4.2.4.2 In-flight Power Supply...... 4-28 4.2.4.3 Optional Services ...... 4-28 4.2.5 Separation Ignition Command ...... 4-28 4.2.6 Payload Telemetry Support ...... 4-28

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List of Figures

Figure 4-1 Multiple payload accommodation for MOM without main payload...... 4-2 Figure 4-2 Integrated MOM payload with main payload simulator...... 4-2 Figure 4-3 Rockot maximum usable payload envelope, all dimensions given in mm.....4-4 Figure 4-4 CASA CRSS Clamp Ring Separation System installation with KSRC pyrolock...... 4-6 Figure 4-5 Stay-out zones (hatched) for the standard interface diameter of 937 mm...... 4-7 Figure 4-6 Stay-out zones (hatched) for the standard interface diameter of 1194 mm...... 4-8 Figure 4-7 Optional CASA LPSS Launcher Payload Separation System, (a) general view, (b) stay-out zones (hatched) for the standard interface diameter of 1194 mm...... 4-9 Figure 4-8 Cut-away detail of the Mechanical Lock System...... 4-10 Figure 4-9 Minisatellite separation system...... 4-11 Figure 4-10 CASA 937 SRF clamp band cylindrical payload adapter. (a) Side view. (b) Launch vehicle interface ring detail...... 4-14 Figure 4-11 CASA CRSS 1194 SRF clamp band conical payload adapter system (Kompsat-2). (a) Side view. (b) Launch vehicle interface ring, detail...... 4-15 Figure 4-12 Elongated conical payload adapter for the CASA CRSS 937 SRF clamp band...... 4-16 Figure 4-13 Example of base-mounted multiple satellite dispenser system for two spacecraft (SMOS/Proba-2)...... 4-16 Figure 4-14 Example of an MLS adapter system for single satellite accommodation. ...4-18 Figure 4-15 Example of side-mounted multiple satellite dispenser system for two spacecraft (GRACE), only one spacecraft attached...... 4-19 Figure 4-16 Example of side-mounted multiple satellite dispenser system for three spacecraft (SWARM)...... 4-19 Figure 4-17 Typical example of an umbilical connector bracket used in combination with a 1194 mm clamp band...... 4-21 Figure 4-18 Umbilical connector OSRS50BATV...... 4-21 Figure 4-19 Rockot umbilical harness diagram...... 4-23 Figure 4-20 Launch site ground wiring diagram...... 4-25

List of Tables

Table 4-1 Pin allocation of umbilical connectors...... 4-23 Table 4-2 Ground wiring capacities...... 4-24 Table 4-3 Bonding / grounding schematic drawing; example of a dispenser configuration...... 4-27 Table 4-4 Operational parameters of the telemetry system TA1...... 4-29 Table 4-5 Operational parameters of the telemetry system TA2...... 4-29

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4 Spacecraft • Dispenser (system): adapters that are built for multiple satellite accommoda- Interfaces tion or side mounted accommodation

Launch vehicle interface ring: the top 4.1 Mechanical Interfaces • part of the payload adapter made of 4.1.1 Payload Accommodation aluminium alloy, part of the adapter interfaces to the spacecraft interface ring The main mechanical interfaces of the and connected to it via a clamp band Rockot launch vehicle to the Customer's separation system spacecraft are described in this section. Examples and illustrations are provided for • Spacecraft interface ring: the ring at- the mounting of single or multiple space- tached to the normally lower part of the craft to the launch vehicle adapter or dis- satellite that interfaces with the adapter penser assembly equipped with a suitable or dispenser system of the launch ve- separation system. Additionally, the allow- hicle, used with clamp band systems able envelope within the Rockot payload • Spacecraft interface points: the inter- fairing is described. face points that are attached to the The following terminology is used within satellite to interface with the adapter or this chapter and defined below to avoid dispenser system of the launch vehicle, confusion: used with the Mechanical Lock System for point attachment • (Payload) adapter: the mechanical structure mounted on the launch vehi- It should be noted that it is standard prac- cle which supports the mated space- tice for EUROCKOT to provide the appro- craft via the spacecraft interface ring priate qualified adapter or dispenser system including all the associated equipment • (Payload) adapter system: the adapter necessary such as a separation system, plus the appropriate hardware to inter- spring pushers, umbilicals and separation face to the spacecraft interface ring or monitoring switches to the customer as interface points, i. e. separation sys- part of the launch services contract. tem, spring pushers, umbilical connec- Hence, the interface of the adapter or dis- tors and separation monitoring penser to the Breeze-KM upper stage is switches entirely the responsibility of EUROCKOT and, therefore, not covered here. • Separation system: the complete separation system including the pyrotechni- To accommodate individual Customers' cal devices and their electrical initiation different needs and satellite designs, system, e. g. a CASA CRSS clamp EUROCKOT offers a wide variety of op- band system or the Mechanical Lock tions for interfacing their spacecraft to the System launcher. The adapter or dispenser systems offered include Russian point attachment separation systems as well as classi-

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cal clamp band separation systems using MainMain Payloadpayload adapterAdapter well established western suppliers. MIMOSA

MOST As well as dedicated single payload SupportSupport Frameframe launches, EUROCKOT also provides diverse accommodation schemes for multiple payloads in order to make maximum use of available resources such as volume and performance.

BreezeBreeze-KM KM As an example, Figure 4-1 depicts a multiple satellite accommodation. This particular arrangement of small satellites around a main payload was exercised for the Multi- ple Orbit Mission (MOM) in June 2003. Eight small satellites weighing from 1 kg to 68 kg accompanying a larger 250 kg main CubeSat "litech" NLS-2 Container payload were accommodated.

The necessary dispensers and adapters for the respective accommodation schemes of the individual satellites will be part of the mission-dependent equipment and will generally be developed and provided by EUROCKOT. Because of the di- NLS-1 Container versity of the mechanical interfaces, only CubeSat UT generic interface details are described here. Further details should be coordinated Figure 4-1 Multiple payload accommodation with EUROCKOT. for MOM without main payload.

Figure 4-2 Integrated MOM payload with main payload simulator.

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4.1.2 Usable Volume for Payload analysis following a coupled loads analysis and will involve the assessment of all The layout of the payload-usable volume, available margins within the envelope. It which is the maximum dynamic envelope should be noted that EUROCKOT recom- above the Breeze-KM interface plane, is mends to all Customers with payload ele- shown in Figure 4-3. ments that are predicted to have less than This figure reflects the maturity of the com- 40 mm clearance from the maximum us- mercial payload fairing design which per- able envelope, e.g. antenna, solar arrays formed its maiden flight in May 2000. The etc., should contact EUROCKOT directly usable volume has been defined conserva- for precise determination of actual clear- tively taking into account the following ances. items: 4.1.3 Spacecraft Accessibility • Maximum dynamic movement of fairing Mechanical access to the payload after • Maximum manufacturing tolerances of encapsulation is not foreseen as a stan- the fully integrated fairing dard service. Principally, fairing access • Maximum mounting error of the fairing hatches are possible provided that an acceptable cleanliness scenario at the • A minimum guaranteed clearance be- Launch Pad can be agreed. However, ac- tween the spacecraft and the fairing cess via umbilical connectors will be provided during any operation phase after • Estimated maximum spacecraft dy- encapsulation, e. g. for battery trickle namic movement for a typical base- charging, communication, etc. Should a mounted configuration. For a side- late intervention at the spacecraft be nec- mounted spacecraft located on a verti- essary, the upper composite will be de- cal payload dispenser, this value will be stacked from the booster unit and trans- the subject of a dedicated dynamic ported back to the payload integration facil- analysis. ity. • Estimated maximum spacecraft mounting error and manufacturing tolerance 4.1.4 Separation Systems of a typical base-mounted adapter system This section describes potential options for providing high quality attachment and EUROCKOT can also provide a three- separation between the Customer’s space- dimensional -IGES- file for a preliminary craft and the Breeze-KM upper stage. The spacecraft accommodation investigation by selection of one of these interface solutions the Customer. is driven by constraints such as spacecraft geometry, mass and related properties, Customers may in certain cases to exceed stiffness and so on. However, cost aspects the maximum dynamic envelope shown and maximum acceptable mechanical above. However, the acceptance of such a loads during spacecraft separation are case is subject to a detailed clearance design drivers, as well.

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A-A B-B C-C

8.56° 16.14° 1 300

120

View Z SEC TION C-C SEC TION B-B SEC TION A-A View Z Breeze-KMBreeze Interface interface A TentativeTentative spaceSpace forfor planePlane with S/C SC Adapter adapter Breeze-KMLV Batteries batteries *) *) C Vertical split line

B B

Dynamic Envelope C Dynamic envelope A *) Potential interference zone with Breeze-KM batteries for extended missions. Please co-ordinate with EUROCKOT for precise information. Please note that the stay-out zone of 300 mm at the Breeze-KM interface plane may be reduced to 100 mm after the successful qualification of new batteries in 2012.

Figure 4-3 Rockot maximum usable payload envelope, all dimensions given in mm.

Three types of separation system which • A minisatellite separation system are used to retain and then release the All adapter concepts and separation sys- satellite are offered: tems described in this chapter, except for • Flight-proven Marmon clamp bands the Launcher Payload Separation System from EADS CASA (LPSS), are flight-proven and can be procured as off-the-shelf equipment. Other • The KSRC-supplied, flight-proven Me- types of separation system can be consid- chanical Lock System

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ered and developed upon Customer's re- ration plane to the spacecraft bottom has quest. The Customer can also provide his to be respected as axial stay-out zone. The own separation system. In this case the figures are provided for example only. For choice of the adapter has to be agreed with formal interface data, EUROCKOT pro- EUROCKOT. vides the dedicated specification ESPE- 0018. The separation system also comprises spring loaded pushers, separation monitor- Taking into account the maximum Rockot ing switches and umbilical connectors at- payload performance, CASA clamp rings tached to suitable brackets which are items are qualified for standard interface diame- to be accommodated in the payload ters of 937 and 1194 mm. The systems are adapter. The pushers can be selected for procured with low shock clamp band separation velocities between 0.1 and 0.8 opener devices which are fully qualified. m/s upon Customer's request. The spring Typical shock response spectra for these pushers are so aligned that the resultant systems are given in chapter 5. For even force will act along the spacecraft’s centre lower separation shock spectra, the EADS of mass in the desired separation direction, CASA Launcher Payload Separation Sys- thereby reducing the spacecraft tip-off rate. tem (LPSS) can adapted for Rockot. For specific information on low shock systems, 4.1.4.1 Clamp Band Separation Systems the Customer is advised to contact The EADS CASA Clamp Ring Separation EUROCKOT directly. System (CRSS) is proposed for use on The choice of the manufacturer for the Rockot payload adapters. The system is adapter as well as specific data on inter- applied with a thermal cycle using electrical face requirements (hole patterns, stay- out heaters. Figure 4-4 to Figure 4-7 provide zones, electrical connectors etc.), are sub- examples of compatible CASA clamp jects for mutual discussion and develop- bands including typical interface and stay- ment, as EUROCKOT offers maximum out zone details. Please note that for the design flexibility to its Customers. CRSS a distance of 60 mm from the sepa-

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UmbilicalUmbilical connector Connector bracket Bracket (2 pieces) (2 Brackets) SCSC interface Interface ring Ring

Catcher

PyrolockPyrolock CRSS CRSSClamp clamp Ring ring (open) (open)

LV interface ring AdapterAdapter LV Interface Ring

Figure 4-4 CASA CRSS Clamp Ring Separation System installation with KSRC pyrolock.

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Figure 4-5 Stay-out zones (hatched) for the standard interface diameter of 937 mm.

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STAYOUT ZONES FOR INSTALLATION AND RELEASE

Figure 4-6 Stay-out zones (hatched) for the standard interface diameter of 1194 mm.

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Bushing Main beam

Pyro-nut

Pyro-nut support Connector support (a)

(b)

Figure 4-7 Optional CASA LPSS Launcher Payload Separation System, (a) general view, (b) stay-out zones (hatched) for the standard interface diameter of 1194 mm.

4.1.4.2 Mechanical Lock Systems The MLS, which is shown in Figure 4-8, has successfully performed more than 25 The Mechanical Lock System (MLS) is in-flight separations, not only with Rockot offered by EUROCKOT for use with space- but also with the Proton launch vehicle. It craft that are attached to the launch vehicle fastens the satellites to the launch vehicle at discrete points rather than via a ring as payload adapter or dispenser via three or in a standard clamp band system. Such four point mechanical attachments at the mechanically driven point attachment sys- base of each satellite. The number of at- tems are particularly advantageous when tachment points depends on the satellite deploying several satellites during a single shape and mass. The spacecraft are re- launch. leased by the firing of a single pyrodriver located in the payload adapter system.

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This actuates a mechanical drive to unlock tenuating the pyrotechnic shock levels ex- the attachment points. The mechanical perienced by the spacecraft. All compo- shock is not exerted directly on the space- nents of the MLS are contained and no craft interfaces, but on the elements of the part will be released. mechanical drive, thereby significantly at-

Separation Plane

Adapter

Note: The indicated parts remain on the spacecraft after separation: 1,7,10 = Bolt, 2,8,9,11 = Washer, 3 = Screw- nut, 4 = Spring, 5 =Support, 6 = Bolt Retainer. Figure 4-8 Cut-away detail of the Mechanical Lock System.

are mounted on the housing of this system. 4.1.4.3 Minisatellite Separation System The upper ring of the separation system is Eurockot also offers a KSRC separation attached to the spacecraft lower bottom by system for minisatellites. An overview and bolts. The spring pushers interface either design details are shown in Figure 4-9. All directly with the spacecraft structure or with the components of the separation system the spacecraft adapter.

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Figure 4-9 Minisatellite separation system.

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4.1.5 Payload Adapters The upper part of the adapter allows for the accommodation of electrical connectors to Payload adapter and dispenser systems, the spacecraft via support brackets. The which are the structures supporting the bracket position can be varied on a case- chosen separation system, are described by-case basis, each bracket also allows in this section. ± 4 mm horizontal and ± 2 mm vertical ad- The payload adapter also accommodates justment for fine tuning. The lower part of the spring loaded pushers, separation the adapter allows for positioning of sepa- monitoring switches, and umbilical connec- ration system components and sensors. tors which are components of the separa- Figure 4-10 depicts a flight qualified tion system (section 4.1.4). cylindrical aluminium payload adapter with 400 mm height together with interface de- 4.1.5.1 Clamp Band Separation System tails. This cylindrical adapter type is cur- Adapters rently in modification for the 1194 mm clamp band. Figure 4-11 shows a conical Adapter systems compatible with classical shaped adapter with an 1194 mm interface Marmon-type V-shaped clamp band sepa- ring diameter specially developed for the ration systems are offered by EUROCKOT. Kompsat-2 spacecraft interface. In the ma- Such payload adapter types are flight- jority of cases the payload adapter consists proven and can be offered in two sizes, of two parts, namely an upper part and a 937 mm and 1194 mm. As an option, also lower part bolted together. The upper part other clamp band sizes can be adapted for provides the launch vehicle interface ring Rockot/Breeze-KM if required. and is manufactured from aluminium alloy. The payload adapters have the shape of The lower part of the adapter which inter- either a cone or a cylinder and are offered faces to the launch vehicle can be made as either aluminium or carbon fibre struc- from a carbon fibre or aluminium structure. tures or a combination of both. The pay- A distinctive feature of these adapter de- load adapter is bolted to the top of the signs is that, while the height of the lower equipment bay of the Breeze-KM upper part may vary for different payloads, the stage. The forward face of the payload payload adapter interface remains un- adapter is machined into a ring with pre- changed. Finally, Figure 4-12 shows a defined dimensions according to the re- conical payload adapter with an extended quirements of the clamp band separation length for the CRSS 937 SRF clamp band. system chosen. The payload adapter may Two spacecraft on top of each other can also be split into a lower part which inter- be launched on Rockot/Breeze-KM using faces to the upper stage and the LV inter- Multi-Satellite Dispenser systems (MSD). face ring that interfaces the clamp band An example of two base mounted satellites and the spacecraft. The spacecraft inter- with clamp band separation systems de- face ring is pressed on top of this surface signed for the SMOS/Proba-2 mission against to the adapter by adequate ten- (2009) is shown in Figure 4-13. The sioning of the clamp band. smaller spacecraft is transported under- neath the conical payload adapter of the

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larger spacecraft and attached with the Espacio that the cross section of the minisatellite separation system (Figure ring is still enough to ensure the correct 4-9). function of the interface. Therefore, the Customer shall provide the spacecraft The spacecraft ring interfacing with the design concept to EUROCKOT for veri- clamp band and adapter on the launch fication before releasing the design to vehicle side is generally designed by the manufacturing. Details for the EADS Customer. While the spacecraft ring inter- CASA CRSS 937 and 1194 SRF clamp face dimensions are fix, the cross section band separation system, the launch is selected by the Customer, who is re- vehicle adapter and the spacecraft in- sponsible for the structural integrity of the terface ring are given in the latest issue spacecraft ring. EUROCKOT will check of EUROCKOT specification ESPE- together with KSRC and EADS CASA 0018.

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∅945.26Ø945.26 Плоскость разделения ВерхнееAdapter кольцо system ПС Separation plane Upper ringupper of the ring AS Separation plane 100 100

ановки 400±0.5 400±0,5 Support forт electricеля umbilical connectoalataion ofr l

I Каркас Frame Ø10.5H12∅10,5H12 Frame 10 (a) JJ

+0.15 0.05 F O945.26Ø945.26 -0.00 0.1 Å 0.3 Å O940 -0.2 5.84±0.076 1.53±0.03 O916±0,5Ø916±0.5 D CoatingCoating: chemical chemical oxidation oxidation ( D O884±0,5Ø884±0.5 F R1 1 1.53±0.03 5.84±0.076 1.6

15° -0.25°-0,25° 6 6 13 13 R1,5±0,5R1.5±0.5 1.6 R1.27 R3 CoatingCoating: chemicalchemical oxidation oxidation ( 6

ViewD D O930±0.2

2,652.65-0,1-0.1 2,52.5 ViewJ J 0.2x45° 0,2x45° vertical notch Vertical notch R0.5±0.13R0,5±0,13 0.5

R0.25 0.5 0.5

(b)

Figure 4-10 CASA 937 SRF clamp band cylindrical payload adapter. (a) Side view. (b) Launch vehicle interface ring detail.

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Плоскость разделения Separation plane КА II SС

Кронштейны отрывных электросоединителей Support for installation of electric umbilical connector

Кольцо ПС Алюминиевый сплав. AS ring Каркас Casing

II Фитинги стыковки с РБ. Mating bracket (US)

(a)

J J

(b)

Figure 4-11 CASA CRSS 1194 SRF clamp band conical payload adapter system (Kompsat-2). (a) Side view. (b) Launch vehicle interface ring, detail.

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Figure 4-12 Elongated conical payload adapter for the CASA CRSS 937 SRF clamp band.

Figure 4-13 Example of base-mounted multiple satellite dispenser system for two spacecraft (SMOS/Proba-2).

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chanical locks using four or three brackets 4.1.5.2 Mechanical Lock System Payload Adapters at the base of each satellite. The number of attachment points depends on the satel- Payload adapters designed for use with the lite shape and mass. Separation is MLS separation system are individually achieved by igniting the pyro-actuator configured to suit the particular spacecraft which in turn rotates the mechanical locks design. This customisation allows ex- via the mechanism rods releasing the tremely lightweight and low height adapters spacecraft adapter frame. The spacecraft to be realized. EUROCKOT and KSRC is then pushed away by spring pushers. have designed, manufactured and suc- Shock is not exerted directly on the space- cessfully flown over two dozen systems of craft interfaces but on the parts of the me- this kind for commercial Customers. These chanical drive, thus significantly attenuat- adapters are also well suited for multiple ing the pyrotechnic shock levels at the satellite accommodation. spacecraft.

The mass and height of the adapter de- EUROCKOT is able to provide customised pend on the final arrangement, i.e. dimen- payload adapter and dispenser systems for sions of the spacecraft and number of at- multi satellite deployment upon Customer tachment points. Spacecraft interface request. EUROCKOT and its parent com- brackets that stay on the spacecraft after pany KSRC have significant experience in separation will have a mass of approxi- the design, manufacture and qualification mately 2 to 10 kg depending on mass and of these systems. Figure 4-15 shows an geometry of the spacecraft (Figure 4-8). example of a side-mounted multiple satel- Adapter heights as low as 100 mm can be lite dispenser. This particular system was realized. designed and qualified for the NASA-DLR GRACE mission. The riveted aluminium An example of a single satellite adapter structure incorporates the mechanical lock using the mechanical lock system is shown system described and allows two space- in Figure 4-14. It incorporates the me- craft to be accommodated side-mounted. chanical lock separation components in- Figure 4-16 depicts the multiple satellite cluding pyro-actuator, rods, spring push- dispenser developed for the ESA SWARM ers, connectors and bonding provisions. mission with the capability to attach and The satellite is fastened to the launch vehi- release three base-mounted spacecraft. cle payload adapter via four or three me-

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Rod (4 places) Pusher (4 places) Pyroactuator Lock (4 places)

Spacecraft Adapter Frame Simulator

Connector Bracket (4 places) Adapter Frame

PusherPusher bracketBracke t (4(4 places) places)

MLSLock (4 places)places) Rod (4 places)

Pyroactuator

Figure 4-14 Example of an MLS adapter system for single satellite accommodation.

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Figure 4-15 Example of side-mounted multiple satellite dispenser system for two spacecraft (GRACE), only one spacecraft attached.

SC push away mechanism SC fastening mechanism +Zsc

III IV +Ysc

+Ysc

+Zsc

II I +Zsc +Ysc

Adapter system

Figure 4-16 Example of side-mounted multiple satellite dispenser system for three spacecraft (SWARM)

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4.2 Electrical Interfaces dynamics, alternative connector types have to be mutually agreed with EUROCKOT. This section describes the interfaces employed to provide electrical links between 4.2.1.2 Separation Verification the spacecraft’s umbilical connectors and the Customer’s EGSE for spacecraft use. The spacecraft manufacturer has to pro- The electrical interfaces include the launch vide spacecraft separation monitoring cir- vehicle to spacecraft on-board electrical cuits as jumpers in each of the spacecraft interfaces, EGSE interfaces, and teleme- umbilical connectors for use by the launch try/command links. vehicle telemetry. Separation monitoring circuits as jumpers on Launch Vehicle side 4.2.1 On-board Interfaces are provided to be used by Spacecraft telemetry. The on-board electrical interfaces provide links from the spacecraft’s umbilical con- 4.2.1.3 Interface Electrical Constraints nectors via the ground cable network to the spacecraft EGSE, including power circuits, The following restrictions on the spacecraft checkout and control circuits, and teleme- to launch vehicle electrical interfaces will try and command links. Examples of the apply: umbilical connector brackets that can usu- • The maximum voltage on the space- ally be accommodated on the payload craft umbilical connectors must not ex- adapter are illustrated in Figure 4-17 and ceed 100 V. At lift-off, the transit cable Figure 4-18. The brackets may vary in de- must be de-energised both on the sign depending on their location on the spacecraft and EGSE side, except for spacecraft and electrical connector types. the separation jumpers.

4.2.1.1 Umbilical Connectors • The EGSE provided by the spacecraft contractor must be designed to inhibit As a baseline EUROCKOT proposes the voltages above 100 V. use of four 50-pin Russian connectors type OS RS50BATV. The connectors are • GSE power through the spacecraft mounted on the payload adapter by means umbilical connectors must be switched of suitable brackets. Wires up to AWG 22 off automatically if the nominal operat- size with a cross section of 0.35 mm2 can ing current is exceeded by 50% over a be accommodated by the OS RS50BATV 0.2 s period. The EGSE supplied by the connectors. The maximum permissible Customer must be designed so that it cross section area of the accommodated will cause automatic switch-off if the wire is 0.5 mm2 on the condition that it is nominal operating current of the space- used via a pin and not more than 20 pins craft lines exceeds 100% over a 0.1 s are used on the contact region connector period. perimeter. The number of connectors and number of pins per connector can be se- • Not later than 20 s before spacecraft separation the current at spacecraft to lected according to the Customer's needs. Hence, the spacecraft connector configura- launch vehicle interface should be lim- tion shown in is only an example. Because ited to 100 mA, except for the separation jumpers. of the potential impacts on the separation

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BB C-C „K„@ DBAS connectorDBAS Connector SeparationSeparation planeplane S„R ViewB B

LVLV axisaxis ±2±2„}„} P

Support ±4„}„} III(I) ±4 ±4„}„} +Xsc(-Xsc) XSC ViewP P „P„R „@S BracketBracket

**Dimension Dimension for for referential referential use use

Figure 4-17 Typical example of an umbilical connector bracket used in combination with a 1194 mm clamp band.

±2*

±2* Connector axis

±2*

* Dimension for referential use

Figure 4-18 Umbilical connector OSRS50BATV.

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4.2.1.4 Umbilical Harness Configuration and ary column. The wires are symmetrically Specifications distributed between the two umbilical ca- The standard Rockot umbilical harnesses bles. Wires of type MC-15-11-0.35 are is schematically shown in Figure 4-19. used. The maximum operating voltage is Four spacecraft umbilical connectors 801 100 V on the spacecraft umbilical connec- to 804 may accommodate up to 200 transit tors, the maximum operating current is 1.5 lines. Any telemetry channel for the moni- A per transit wire. All transit wires have a 2 toring of the spacecraft interface environ- 0.35 mm cross-section. The total length of ment will share this budget of lines with the the on-board transit lines from connectors spacecraft. These lines lead to the connec- 811 and 813 on the Breeze-KM pressur- tors 811 and 813 of type OS 9R102 ised equipment bay to connectors ShR010 (OC 9P102) with 102 pins each. These and ShR011 on the container plate is less connectors are mounted on the payload than 18 m. adapter / Breeze-KM interface. Telemetry Neglecting the resistance of the payload circuits from the umbilical connectors 801 harness, the resistance of the on-board through 804 are routed into connector 840 cable network from spacecraft connectors and from there to the upper stage teleme- 801 through 804 to the connectors on the try. From connectors 811 and 813 the cir- plate of the launch container ShR010 and cuits are routed into connectors ShR10 ShR011 is not more than 1 Ohm. The and ShR11 with 102 pins each on the shields of the twisted pairs and single Breeze-KM umbilical plate to the ground shielded wires are isolated from the cable interface. The launch container plate ac- jackets and launch vehicle connector shells commodates electrical connectors of type and terminated at the appropriate electrical OS RRM47 with 102 pins each. connector pins. The circuitry of the payload adapter har- The insulation resistance of the transit lines ness from the spacecraft umbilical connec- should be at least 10 MOhm. tors 801 through 804 to connectors 811 and 813 is developed on the basis of the A high reliability is ensured by: Customer’s input data and is payload- • Highly reliable components operated in specific. Table 4-1 shows the pin allocation a derated mode requirements. It is however limited by the capacities of the harness of the Breeze-KM • Verified service life margins as to oper- that is specified in Table 4-2. The harness ating time, storage time and number of length from the spacecraft umbilical con- actuations nectors 801 to 804 to connectors 811 and 813 depends on the payload adapter de- • Verified robustness and environmental sign. resistance margins

The harness beyond the connectors 811 and 813 down to the spacecraft EGSE in Undertable Room 7 is standardised transit wiring via the upper stage and the station-

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J1 J2 J3 J4 SC Separation plane 50 80150 80250 80350 804

LV AS

T840 Telemetry system 811 813 102 102 Breeze-KM pressurized 2ШРВ2 2ШРВ3 instrument compartment 2ShRV2 2ShRV3 102 102 Breeze-KM intermediate compartment ШР10 ШР11 ShR10 ShR11 US separation circuit board 102 102

Launch Container ШР010 ШР011 ShR010 ShR011 102 102

Figure 4-19 Rockot umbilical harness diagram.

Signal Max. Max. Max. Specific Pin Line start Powered designation voltage, current, Resistance, Line end require- No. (Source) or signal or type V А Ohm ments

example of table to be filled in by the Customer during the mission integration process

1. In the column “Specific requirements” please specify: x Single, non shielded wire twisted shielded pair. x With what umbilical electrical connector pin does it combine to make twisted shielded pair? x For jumpers on the spacecraft or launcher side side, with what umbilical electrical connector pin does it combine to make a jumper? 2. The resistance values specified by the Customer apply from the spacecraft to umbilical connectors through the Customer EGSE in Undertable Room 7. Table 4-1 Pin allocation of umbilical connectors.

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Total transit 2 Transit wire type Quantity Wire cross section, mm Note wires Flight harness Ground harness Single, no shield 100 100 0.35 2.5

Single shielded 30 30 0.35 1.5

Twisted shielded pairs 34 68 0.35 2.5

Shield 2 2 0.35 >1.5 Total: 200

Table 4-2 Ground wiring capacities.

4.2.1.5 Matchmate / Electrical Checkout erations involving the spacecraft EGSE, as well as upper composite integration and The Matchmate and electrical checkout mating with the launch vehicle at the proc- between the spacecraft interface and the essing facility. payload adapter is strongly recommended and should preferably be conducted at the Serving as an electrical extension of the spacecraft manufacturer's facility. launch vehicle on-board harness ( and Figure 4-20), the ground wiring is consis- 4.2.1.6 Spacecraft Electrical Interface Input tent with all electrical characteristics as Data Requirements applicable to the spacecraft on-board equipment lines. The ground wiring length For the purpose of Rockot mission adapta- is approximately 65 m from electrical con- tion, the Customer shall provide input data nectors ShR010 and ShR011 to electrical containing specifications for each transit connectors X1-1, X2-1, X3-1, X4-1. The wire line per umbilical connector, as shown total number of ground wires to support the in Table 4-1. spacecraft on-board equipment and the spacecraft EGSE is 200. The configuration 4.2.2 Ground Electrical Interface of the ground harness is identical to the The ground wiring is designed to interface harness installed on the Breeze-KM (Table between the launch vehicle on-board har- 4-2). ness and the spacecraft EGSE located in The ground wiring can be terminated with Undertable Room 7. any electrical connector as required for the The ground wiring will only be used to sup- spacecraft EGSE interface in Under-table port payload electrical testing or other op- Room 7.

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Fairing

S/CSC

Stationary mast ШР011 ShR011 S2C2 Stationary Mast ShR011

ШР010ShR010 S1C1 TransportTransport andand UpperUpper ShR010 launchLaunch container Container Stagestage extensionExtension

X3- ST X1- ST X2- ST X4- ST X1-1 X2-1 X3-1 X4-1 Undertable 50 50 50 50 Room TransportTransport and launchLaunch Plate containerContainer

Plate

X1-3 X2-3 X3-3 X4-3 Blockhouse (Optional) LV

Figure 4-20 Launch site ground wiring diagram.

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4.2.3 Payload Grounding and Bonding adapter system should be as defined below: A passive approach is employed to protect the upper composite from hazardous static • Alodine 1200 conductive coating on the build-up. This approach includes bonding, spacecraft side creation of conductive surfaces, and • SECO electro conductive oxidizing on grounding of the upper composite. The the adapter system side goal of the three techniques is • The transient resistance should not • to reduce the voltage potential between exceed 10 mOhm. any two structural elements to a safe level, The upper composite ESD control is implemented by: • to ensure more effective cable shield- ing in order to eliminate any electro- • the use of external surface materials static discharge (ESD) risk for the per- with a volume resistivity of less than sonnel, and 105 Ohm · m,

• to reduce the voltage difference be- • coating non-conductive materials with tween the upper composite or any of its conductive layers to be bonded to the components and the ground to zero. metal structure,

The upper composite is designed • the use of a conductive film, foil, grid or fabric to create a conductive outer sur- • to have no outer surface areas with face in a dielectric, voltage drops equal to the threshold beyond which a hazardous electrostatic • bonding each spacecraft to the dis- discharge becomes possible, penser/adapter by means of two umbilical straps, • to ensure reliable electrical bonding of all metal elements of the structure so • bonding any upper composite compo- that a common reference, or a common nent with at least two points separated electrical mass will be created, by the maximum possible distance, and

• to ensure that any charge that may • electrically interconnecting all layers of build up on an outer conductive surface each multi layer insulation blanket by of a dielectric component will leak a bonding each blanket to the metal way to the common reference, and structure.

• to enable grounding of the upper com- To prevent ESD, shield braiding and transit posite during integration, testing, fuel- cable connector bodies should be con- ling, and transportation. nected to the launch vehicle.

The usage of conductive coatings on the On the pad, the upper composite is mating surfaces of the spacecraft and the grounded via the launch vehicle metal structure. For this purpose, the spacecraft

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is connected to the adapter or dispenser A grounding point is envisaged at the up- via bonding straps or a conductive coating per composite for grounding the upper applied on the spacecraft and adapter mat- composite in the course of manufacturing, ing surfaces. The adapter system in turn is processing, handling and transportation. connected to the upper stage / launch ve- The resistance across the interface at any hicle via bonding straps. Two detachable bonding or grounding point must not ex- bonding straps or a conductive coating is ceed 2 x 10-3 Ohm. applied to the spacecraft and adapter system mating surfaces to ensure adequate Upper composite ESD control is achieved electrical contact between the spacecraft as shown in the bonding / grounding and the dispenser. schematic in Table 4-3.

For this purpose, the spacecraft is required The upper composite will be protected from to have an "earth" reference point close to direct lightning hits by the launch facility the separation plane, on which a bonding lightning protection system. strap can be mounted. The contact resistance at the bonding points is required to be less than 3 x 10-3 Ohm.

2 4 SpacecraftSATELLITE 7 Payload fairing Upper stage

1 3 5 6 Booster

Intermediate section

Recommended bonding Number of Components to be bonded technique bonding points Comments

1 Payload fairing Continuous conductive coating Entire surface 2 Spacecraft/Adapter Detachable straps 2 3 Dispenser/equipment bay Non-detachable straps 2 4 PLF halves/upper stage Detachable straps 2 5 Intermediate section / upper stage Detachable straps 2 6 Upper stage Continuous conductive coating Entire surface Intermediate section / 7 Non-detachable straps 4 booster stack

Table 4-3 Bonding / grounding schematic drawing; example of a dispenser configuration.

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4.2.4 Payload Auxiliary Power Supply terface connector as used for the payload auxiliary power lines. 4.2.4.1 Ground Auxiliary Power Supply The pyrotechnic command for spacecraft An uninterruptible power supply (UPS) is separation is a standard provision. provided for the spacecraft EGSE. Please refer to Chapter 10 for further details. 4.2.6 Payload Telemetry Support

4.2.4.2 In-flight Power Supply The Rockot on-board telemetry system comprises a low rate telemetry device TA1 At launch and in flight up to payload de- and a high rate telemetry device TA2. TA1 ployment, the payload can be supplied with can operate up to end of Breeze-KM op- power from the batteries of the upper stage eration in three modes: as an optional service. The payload can be supplied with: • DT: Direct transmission

• Power (non-stabilised) with 24 to 30 • REC: Data record VDC (voltage spikes of +/- 3 V might be • REP: Data replay encountered within this range), in duration of up to 50 ms REC is used for the flight phases without visibility to downlink the data in the subse- • Maximum power supply 15 Ah for 7 hours with not more than 5 A quent visibility phase in the REP mode. The total storage capacity of TA1 is 64 4.2.4.3 Optional Services Mbit.

For a customer-supplied separation system The following channels in the TA1 system power supply can be provided as an op- are assigned to the payload: tion, with the following characteristics: • 24 event channels with 1 bit • Voltage: 28 VDC • 20 analogue channels with 8 bit resolu- • Current: pulse of 10 A and 30 ms duration tion for up to 10 pulses • 10 temperature channels with 8 bit re- solution 4.2.5 Separation Ignition Command The maximum data sampling and downlink Discrete sequencing commands, gener- data rates depend on each other and on ated by the Rockot on-board computer, are the operating mode as listed in Table 4-4 available to the payload during the payload and Table 4-5. injection phase. TA1 also registers and transmits the status The number of command lines provided for signals of the payload separation. The the payload, as well as the signal charac- separation signals are generated by the LV teristics, will be defined in detail in the In- separation detection devices. terface Control Document. Discrete lines are provided through the same type of in-

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TA2 operates up to the second stage sepa- • Five channels, each with 500 Hz sam- ration in a direct downlink transmission pling rate mode with a maximum data transmission TA1 and TA2 can provide channels for the rate of 320 kBit/s. The following channels data acquisition from the payload adapter are assigned for payload needs: or dispenser re-allocated within the overall • Three channels, each with 8000 Hz limits as specified above. sampling rate

Data Sampling rate, Hz transmission Operating rate, event analogue temperature Time of mode mode kBit/s channels channels channels realization

DT 1 256 50 50 0.4 on LC, in flight DT 2 32 6.25 6.25 0.05 after SC separation REC 1 256 50 50 0.4 in flight REC 2 32 6.25 6.25 0.05 in flight REC 3 4 0.78 0.78 0.006 in flight REP 1 256 as recorded in REC 1-3 in flight REP 2 32 as recorded in REC 1-3 in flight

Table 4-4 Operational parameters of the telemetry system TA1.

Data transmission Operating rate, mode kBit/s Time of mode realization

DT 320 in flight until second stage separation

Table 4-5 Operational parameters of the telemetry system TA2.

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Chapter 5 Spacecraft Environmental Conditions

Table of Contents

5 Spacecraft Environmental Conditions ...... 5-1 5.1 Mechanical Environment ...... 5-1 5.1.1 General ...... 5-1 5.1.2 Quasi-Static Accelerations...... 5-2 5.1.3 Low Frequency Vibration ...... 5-3 5.1.4 Acoustic Noise ...... 5-4 5.1.5 Random Vibration ...... 5-5 5.1.6 Shock...... 5-5 5.1.7 Loads during Ground Operations ...... 5-7 5.1.7.1 Handling Loads During Ground Operations ...... 5-8 5.1.7.2 Spacecraft Container Railway Transportation Loads ...... 5-8 5.1.7.3 Roadway Autonomous Transportation in Archangel and Plesetsk Cosmodrome ...... 5-9 5.1.7.4 Loads during Transport in Upper Composite...... 5-10 5.2 Thermal Environment ...... 5-11 5.2.1 Environmental Conditions in the Integration Facility...... 5-11 5.2.2 Pre-Launch Temperature Control within the Fairing...... 5-11 5.2.3 In-flight Temperature under the Fairing ...... 5-13 5.2.4 Aerothermal Flux at Fairing Jettisoning ...... 5-14 5.2.5 Heat Impact during Coasting Phase ...... 5-14 5.3 Fairing Static Pressure during the Ascent ...... 5-14 5.4 Contamination and Cleanliness...... 5-15 5.5 Electromagnetic Environment ...... 5-15 5.5.1 Launch Vehicle Electro-Magnetic Emissions...... 5-15 5.5.2 Spacecraft Electro-Magnetic Emissions ...... 5-17

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List of Figures

Figure 5-1 Launch vehicle and payload coordinate system...... 5-1 Figure 5-2 Typical axial payload interface acceleration...... 5-2 Figure 5-3 Typical radial payload interface acceleration...... 5-3 Figure 5-4 Low frequency vibration environment at the separation plane...... 5-4 Figure 5-5 Composite noise spectrum under the Rockot payload fairing during flight given in 1/3 octave band referenced to 2 x 10-5 Pa RMS...... 5-4 Figure 5-6 Typical shock environment for diverse separation systems...... 5-6 Figure 5-7 Transportation of the spacecraft in the spacecraft container...... 5-7 Figure 5-8 Spacecraft transportation as part of the upper composite...... 5-7 Figure 5-9 Random vibration power spectrum acting at spacecraft container base during autonomous railway transportation...... 5-8 Figure 5-10 Random vibration spectrum for autonomous roadway transportation...... 9 Figure 5-11 Random vibration power spectrum acting on the spacecraft during upper composite transportation...... 10 Figure 5-12 Thermal conditioning of the upper composite during transportation...... 5-12 Figure 5-13 Thermal conditioning of the upper composite at the launch pad...... 5-12 Figure 5-14 Variation of fairing inner static pressure during ascent...... 5-15 Figure 5-15 Field strength of launch vehicle and cosmodrome electromagnetic emissions in the adapter plane without payload fairing...... 5-16 Figure 5-16 Allowable spacecraft emissions at Plesetsk Cosmodrome...... 5-17

List of Tables

Table 5-1 Quasi-static loads experienced by an average payload during ascent...... 5-3 Table 5-2 Low frequency vibration environment...... 5-3 Table 5-3 Composite noise spectrum under the fairing during flight...... 5-5 Table 5-4 Typical shock environment for diverse separation systems...... 5-6 Table 5-5 Maximum handling loads...... 5-8 Table 5-6 Shock loads during autonomous railway transportation...... 5-8 Table 5-7 Random vibration power spectrum acting at spacecraft container base during autonomous railway transportation...... 5-8 Table 5-8 Quasi-static accelerations for autonomous roadway transportation at wind velocities ≤ 20 m/s...... 5-9 Table 5-9 Random vibration power spectrum for autonomous roadway transportation...... 5-9 Table 5-10 Accelerations during upper composite transportation at wind velocities ≤ 20 m/s...... 5-10 Table 5-11 Random vibration power spectrum acting on the spacecraft during upper composite transportation...... 5-10 Table 5-12 Air conditioning systems performance data...... 5-13 Table 5-13 Ranges of measured parameters under the payload fairing...... 5-13 Table 5-14 Launch vehicle and cosmodrome electromagnetic emissions...... 5-16 Table 5-15 Allowable spacecraft emissions at Plesetsk Cosmodrome...... 5-17

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5 Spacecraft gime, or may be due to loading induced by the operation of the propulsion systems Environmental due to longitudinal acceleration, thrust Conditions build-up or tail-off transients, structure- propulsion coupling, attitude control operation, etc. This section describes the environment to which the spacecraft is exposed during the The various types of mechanical environ- launch site operations and its transport into ment experienced by the payload are orbit with the Rockot launcher as well as described in the following paragraphs. accelerations occurring during ground Typical data are given for sine, random transportation and handling. and shock environments. If not explicitly stated, all mechanical loads are defined as maximum operational loads. They are 5.1 Mechanical Environment estimated with 90% confidence and their levels will not be exceeded in 99% of all 5.1.1 General flights. For different dispenser types, dedi- During flight, the payload is subjected to cated environments have to be defined on static and dynamic loads induced by the a case-by-case basis. launch vehicle. Such excitation may be of For the launch vehicle and payload coordi- aerodynamic origin, especially wind and nate system refer to Figure 5-1. gusts, buffeting in the transonic flight re-

X Y

Z Y0 X0

Figure 5-1 Launch vehicle and payload coordinate system.

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The spacecraft coordinate system without The low-frequency or quasi-static compo- index may be selected with the origin at the nent of dynamic loading, which is the fre- CoM of the spacecraft or in the attachment quency component below 100 Hz is used plane and with axes parallel to the launch as the dimensioning factor when selecting vehicle coordinate system, if the satellite is technical solutions aimed at ensuring clocked orthogonally. The orbital block adequate strengths of the payload and coordinate system with the index "O" is a adapter system structures. useful tool for the definition of the loads Figure 5-2 and Figure 5-3 show typical low- and performance of the coupled loads frequency spacecraft interface accelera- analysis (CLA). Its origin lies in the Breeze- tions during powered flight of stages 1 and KM / payload adapter interface plane. The 2. Payload accelerations during powered Y and Z axes concur with the launch vehi- flight of the upper stage are much smaller cle stabilisation axes III and IV respec- and, thus, of no concern for the structure’s tively. load capability.

5.1.2 Quasi-Static Accelerations Quasi-static loads experienced by the spacecraft are largely determined by the During ascent, the payload will experience spacecraft mass and stiffness properties. dynamic loads generated by The loads given in Table 5-1 can be used • the operating launch vehicle propulsion as preliminary dimensioning inputs to the unit, design of the spacecraft. These loads are the worst-case limit loads that do not • the aerodynamic forces appearing as necessarily apply to any spacecraft. the launch vehicle traverses dense Instead, they are obtained for an average atmospheric layers, and spacecraft and shall be subject to updates by a dedicated CLA for each particular • the operating launch vehicle navigation mission. system.

2 60

0 Axial acceleration, m/s -20 0 50 100 150 200 250 300 Time, s

Figure 5-2 Typical axial payload interface acceleration.

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2 2

-2

-4 Radial acceleration, m/s acceleration, Radial -6 0 50 100 150 200 250 300 Time, s

Figure 5-3 Typical radial payload interface acceleration.

axial, g Seq. Event max min radial, g

1 Lift-off + 3.6 0 ± 0.7

2 Worst-case dynamic pressure qmax + 2.8 + 2.4 ± 0.9

3 Peak stage 1 thrust Pmax + 8.0 + 6.2 ± 0.5

4 Worst-case axial stage 1 structural acceleration nXmax + 8.1 + 6.3 ± 0.5 5 Stage 1 shutdown + 8.1 – 1.5 ± 0.7 6 Stage 2 powered flight + 3.0 0 ± 0.4 7 Upper stage powered flight + 1.6 0 ± 0.4 Note: 1) "+" means compression, "-" means tension 2) Axial and radial loads may exist simultaneously.

Table 5-1 Quasi-static loads experienced by an average payload during ascent.

These values are subject to revision after 5.1.3 Low Frequency Vibration the performance of a CLA with a dedicated The low frequency longitudinal and lateral structural mathematical model of the vibration environment spectra experienced spacecraft. at the spacecraft separation plane are A notching to design loads can be allowed presented in Table 5-2 and Figure 5-4 based on controller data at the adapter respectively. interface to the spacecraft or at the inter- Acceleration, g face to Breeze-KM whichever is more Frequency, critical. Hz Longitudinal Lateral 5 - 10 0.8 0.5 For test approaches and factors please 10 - 20 0.8 - 1.2 0.5 refer to chapter 6. 20 - 40 1.2 - 0.8 0.5 40 - 100 0.8 0.5

Table 5-2 Low frequency vibration environment.

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10 Longitudinal vibrations Lateral vibrations

1 Acceleration, g

0.1 110100

Frequency, Hz

Figure 5-4 Low frequency vibration environment at the separation plane.

5.1.4 Acoustic Noise pressure of 2 x 10-5 Pa = 0 dB. Noise is substantially lower outside these periods. The spacecraft is exposed to an acoustic environment throughout the boost phase of The composite noise spectrum at lift-off is flight until the vehicle is out of the atmos- given in Figure 5-5 and Table 5-3. The phere. Acoustic noise is generated by given spectra reflect the guaranteed upper engine noise, buffeting and boundary layer limit, including the worst case fill factor for noise. The level is highest at lift-off (137.9 spacecraft within the payload fairing. The dB OASPL) and in the transonic phase composite duration is 10 seconds. (135 dB OASPL) with a RMS reference

135

130

125

120

115

SPL, dB dB SPL, 110

105

100 10 100 1000 10000

Frequency, Hz

Figure 5-5 Composite noise spectrum under the Rockot payload fairing during flight given in 1/3 octave band referenced to 2 x 10-5 Pa RMS.

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1/3 octave band random vibration depends on specific centre frequency, Hz Sound pressure level, dB payload configuration, e.g. fill factor, pay- 25 120.7 load adapter, payload surfaces. Therefore 31.5 123.3 a generic random vibration environment is 40 123.5 not defined here. In the case of small 50 124.2 compact satellites it may be more conven- 63.5 125.0 80 127.4 ient to perform a random vibration test. In 100 128.7 this particular case, the customer is asked 125 127.4 to contact EUROCKOT directly for defini- 160 126.8 tion of the appropriate random vibration 200 130.4 test levels. 250 126.7 315 124.7 5.1.6 Shock 400 122.8 500 124.4 The spacecraft is subjected to a shock 630 121.5 environment during separation of the fair- 800 118.6 ing and launch vehicle components as well 1000 120.7 as spacecraft separation. 1250 114.3 1600 111.1 The shock levels at the spacecraft separa- 2000 111.1 tion plane experienced during the separa- 2500 114.4 tion from the Breeze-KM upper stage are 3150 112.4 associated with the separation system 4000 111.4 selected (section 4.2). 5000 110.4 6300 108.9 The shock levels are also dependent on 8000 107.4 the pre-tension of the bolts for the MLS 10000 106.9 system or the pre-tension of the band for OASPL 137.9 the clamp band systems. Pre-tensions are Note: determined according to the spacecraft • The acoustic loads are referenced to the RMS acoustic pressure of 2x10-5 Pa and interface characteristics such as • Test duration and factors are defined in chapter 6 spacecraft CoM and interface diameter in • Composite duration is 10 seconds relation to the separation system following Table 5-3 Composite noise spectrum under the fairing during flight. qualified standards. Table 5-4 and Figure 5-6 show the shock specifications for the fairing separation as well as the MLS and

5.1.5 Random Vibration EADS CASA clamp band systems with typical pre-tensions. Random vibration is mainly generated by the acoustic noise field under the payload For detailed information for other sizes, fairing and is also transmitted via the special pre-tensions and low shock sys- launcher structure. Random accelerations tems, please contact EUROCKOT. excited by the launch vehicle engines can be neglected for spacecraft design. The Shock loads are defined as worst value applicable to all three axes and they act

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simultaneously. For clamp band systems note that the second stage separation the shock loads are measured in radial, shock levels are lower than the specifica- tangential directions to the clamp band and tions given here, thus not to be considered normal to the separation plane. Please in the shock environment.

Acceleration, g (SRS, Q=10) Small Mechanical satellite Fairing Lock CASA CRSS CASA CRSS separation Frequency, Hz Separation** System* 937 SRF* 1194 system 100 50 50 80 80 40 700 700 800 1000 1000 400 1000 1000 2000 1800 1800 600 1500 1000 2000 2800 2800 2000 1000 2800 2800 1500 4000 1000 4000 2800 2800 5000 1000 4000 6000 1000 1800 10000 1000 2000 1800 * Shock values for separation systems are measured 40 mm above the separation plane. ** Fairing separation shock is defined at the base of the payload adapter. Hence, for taller payload adapters shock at the spacecraft to adapter interface plane will be lower. Table 5-4 Typical shock environment for diverse separation systems.

10000

1000

Acceleration, g Acceleration, 100 Fairing Separation Mechanical Lock System CASA CRSS 937, 1194 Small Satellite Separation System 10 100 1000 10000 Frequency, Hz

Figure 5-6 Typical shock environment for diverse separation systems.

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spacecraft container. Section 5.1.7.2 con- 5.1.7 Loads during Ground Operations tains the loads given for spacecraft container transportation with a railway wagon This section presents the ground loads the as per the co-ordinate definitions in Figure spacecraft or its container is subjected to 5-7. Similarly, section 5.1.7.3 provides on the EUROCKOT standard ground loads for short duration truck transport from transportation route from the port-of-entry the Archangel airport to the railhead in Archangel Talagi Airport to Plesetsk Cos- Plesetsk and within the Cosmodrome, if modrome by lorry and railway. The right- applicable. Finally, section 5.1.7.4 provides handed co-ordinate systems used for the the loads on the spacecraft during railway different ground operations are defined in transportation to the launch pad as part of Figure 5-7and Figure 5-8. All loads that are the Rockot launch vehicle upper compo- specified in this section are 3 sigma values site. The specific measurement plan and with 90% confidence. The accelerations positions of the instrumentation (Figure during the transportation are defined for 5-7, Figure 5-8) can be customised for wind velocities less than 20 m/s. each customer's individual project. Section 5.1.7.1 provides the maximum handling loads to be expected on the

SC container Three-Axis- Accelerometer Direction of Transportation car transportation

Z X

Figure 5-7 Transportation of the spacecraft in the spacecraft container.

Direction of transportation

Z X

Three-Axis- Y Accelerometer 1 Transportation Unit Three-Axis- Accelerometer 2

Figure 5-8 Spacecraft transportation as part of the upper composite.

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5.1.7.1 Handling Loads During Ground Amplitude, Impulse Number Direction Operations g duration, s of loads

The maximum handling loads acting at the X 1.0 spacecraft CoM during the ground opera- Y 1.0 0.16 - 0.035 100 tions, e. g. hoisting, are presented in Table Z 0.5 5-5. While the gravity acceleration acts Note: The impulse form can either be a triangle or half-sine. continuously in the vertical direction, other operational loads may act simultaneously Table 5-6 Shock loads during autonomous railway transportation. along the other directions.

Quasi-static loads Fre- Power spectral density, g²/Hz quency, direction Vertical Lateral Hz direction directions Safety factor X Y Z 1±0.5 ±0.3 1.5 2 1.5E-4 7.5E-4 6.0E-4

Table 5-5 Maximum handling loads. 4 8.0E-4 1.0E-2 8.0E-4 8 3.0E-3 1.0E-2 1.0E-3

5.1.7.2 Spacecraft Container Railway 10 1.0E-3 1.0E-2 3.0E-3 Transportation Loads 14 8.0E-4 3.0E-3 1.0E-3 20 8.0E-4 1.0E-3 1.0E-3 The location of the accelerometers and the co-ordinate system for the load measure- 25 8.0E-4 8.0E-4 1.0E-3 ment during the spacecraft container trans- 30 8.0E-4 1.5E-3 1.0E-3 port by railway are defined in Figure 5-7. 35 1.2E-3 8.0E-4 1.0E-3 The loads for railway transportation from 40 4.0E-4 6.0E-4 1,5E-4 Archangel to Plesetsk are defined in Table 45 4.0E-4 4.3E-4 1.5E-4 5-6, Table 5-7 and Figure 5-9. 50 4.0E-4 2.8E-4 1.5E-4 Time, min 420 420 420 Table 5-7 Random vibration power spectrum acting at spacecraft container base during autonomous railway transportation.

10-1 10-1 10 -1

10-2 10-2 10 -2 /Hz /Hz /Hz 2 10-3 2 10-3 2 10 -3 , g , g , g Z X Y S S S 10-4 10-4 10 -4

10-5 10-5 10 -5 0 10 20 30 40 50 0 1020304050 0 1020304050 Frequency, Hz Frequency, Hz Frequency, Hz

Figure 5-9 Random vibration power spectrum acting at spacecraft container base during autonomous railway transportation.

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5.1.7.3 Roadway Autonomous Transportation in Archangel and Plesetsk Cosmodrome Power spectral density, g2 Hz Frequency, direction The quasi-static acceleration loads for the Hz X Y Z short duration roadway transportation in 2 5.0E-5 7.0E-4 4.0E-4 Archangel are defined Table 5-8, and 4 1.2E-4 3.0E-3 5.5E-4 Figure 5-10. The random vibration loads 8 3.0E-4 3.0E-3 7.0E-4 are defined in Table 5-9. These spectra 10 3.0E-4 3.0E-3 7.0E-4 apply also to the short road transportation 14 3.0E-4 3.0E-3 7.0E-4 at the launch site in the Plesetsk Cos- 20 1.0E-4 4.0E-4 3.0E-4 modrome. The reference coordinate sys- 25 4.0E-5 1.5E-4 1.0E-4 tem definition given in Figure 5-7 also 30 1,3E-5 5.0E-5 4,5E-5 applies to the roadway autonomous trans- 35 6.0E-6 2.0E-5 2.0E-5 40 3.0E-6 1.0E-5 1.0E-5 portation. 45 2.0E-6 6.0E-6 6.0E-6 Quasi-static accelerations, g Safety 50 1.0E-6 3.2E-6 3.2E-5 X Y Z factor Time, min 10 10 10 ±1.0 1±1.0 ±0.5 1.5 Table 5-9 Random vibration power spec- Note: For transportation regimes the loads are trum for autonomous roadway applied according to the coordinate system of the transportation. transportation system. The loads may apply at the same time on axes X, Y and Z. Table 5-8 Quasi-static accelerations for autonomous roadway transportation at wind velocities ≤ 20 m/s.

10-2 10-2 10-2

10-3 10-3 10-3 /Hz /Hz /Hz 2 10-4 2 10-4 2 10-4 , g , g , g Z X Y S S S 10-5 10-5 10-5

10-6 10-6 10-6 0 10 20 30 40 50 0 1020304050 01020304050 Frequency, Hz Frequency, Hz Frequency, Hz

Figure 5-10 Random vibration spectrum for autonomous roadway transportation.

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Power spectral density, g2 /Hz 5.1.7.4 Loads during Transport in Upper Frequency, direction Composite Hz X Y Z The location of the accelerometers and the 2 6.0E-5 6.0E-5 1.2E-4 co-ordinate system for spacecraft transport 4 2.0E-4 1.0E-4 1.5E-4 within the upper composite are defined in 8 4.5E-4 5.0E-4 2.0E-4 10 1.5E-4 5.0E-4 3.0E-4 Figure 5-8. The upper composite is trans- 14 1.2E-4 5.0E-4 1.0E-4 ported over a distance of 7 km at 3 to 5 20 1.2E-4 1.5E-4 1.0E-4 km/h. The corresponding loads are defined 25 1.2E-4 1.2E-4 1.0E-4 in Table 5-10, Table 5-11 and Figure 5-11. 30 1.2E-4 1.5E-4 1.0E-4 The loads may act simultaneously in the 35 2.0E-4 1.1E-4 1.0E-4 respective axes. 40 1.5E-4 1.0E-4 3.7E-5

45 1.0E-4 8.3E-5 3.7E-5 Quasi-static accelerations ,g Safety 50 1.0E-4 7.5E-5 3.7E-5 X Y Z factor Time, min 18 18 18

±1.0 1±0.5 ±0.3 1.5 Table 5-11 Random vibration power spectrum acting on the spacecraft dur- Table 5-10 Accelerations during upper coming upper composite transporta- posite transportation at wind ve- tion locities ≤ 20 m/s.

10-3 10-3 10-3

10-4 10-4 10-4 /Hz /Hz /Hz 2 2 2 , g , g , g Z X Y S S 10-5 S 10-5 10-5

10-6 10-6 10-6 0 10 20 30 40 50 01020304050 0 10 20 30 40 50 Frequency, Hz Frequency, Hz Frequency, Hz

Figure 5-11 Random vibration power spectrum acting on the spacecraft during upper composite transportation.

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5.2 Thermal Environment 5.2.2 Pre-Launch Temperature Control within the Fairing This section describes the thermal envi- After encapsulation and upper composite ronment of the spacecraft before and after integration, conditioned air to the fairing is launch. For the definition of the spacecraft supplied either by a mobile thermal condi- thermal environment, three phases of the tioning system on a railcar or a stationary mission are considered: thermal conditioning system at the launch • The spacecraft preparation phase pad. A removable thermal cover is installed within the preparation buildings, on the outer fairing surface for the duration of the upper composite transportation from • transportation to the launch pad, when the integration facility to the launch site, the the spacecraft is encapsulated inside upper composite lifting into the service the fairing, and tower, and the upper composite installation • the in-flight environment phase after on the booster unit. A supplementary mating the upper composite with the spacecraft battery air conditioning system launch vehicle. can be provided as an optional service. The upper composite air conditioning 5.2.1 Environmental Conditions in the configuration during transportation from the Integration Facility integration facility to the launch pad and The payload is processed in clean rooms while at the launch pad is shown in Figure of the integration facility MIK. The clean 5-12 and Figure 5-13. rooms are compliant with ISO Standard The basic performance data of the mobile 14644-1 Class 8 (former FED-STD-209 air conditioning system, the characteristics Class 100,000) with a regulated tempera- of the air supplied by the upper composite ture of 18 to 25 °C and relative humidity stationary air conditioning system, and the between 30 and 60%. EUROCKOT can characteristics of the air supplied by the also provide ISO Class 7 cleanliness (for- optional spacecraft battery air conditioning mer Class 10,000) as an optional service. system are shown in Table 5-12.

The mobile and stationary air conditioning systems are compatible with ISO Class 8 cleanliness, ISO Class 7 can be provided as an optional service. The mean air velocity induced by the thermal conditioning system within the fairing does not exceed 3 m/s.

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Fairing with inner Upper composite Upper Composite thermal insulation

AAirir ducts Ducts ThermalMobile thermal Conditioning conditioning Unit unit

ThermalThermal Cover cover

Transportation Transportationunit Unit

Figure 5-12 Thermal conditioning of the upper composite during transportation.

Temperature and relative humidity sensors in the payload fairing AirAir Ductduct Temperature and relative humidity Umbilical Connect Umbilical connector sensors at the payload adapter

Stationary Airthermal Inlet conditioning(Ø 200mm) unit

Channel (5Channel Places) (5 places) StationarStationaryy column Mast Air Outlet (4AirPlaces) outlet (4 places)

Figure 5-13 Thermal conditioning of the upper composite at the launch pad.

The air temperature inside the fairing is metric mathematical models to be provided between 10 and 25°C during active tem- by the Customer. The adapter hardware perature control, and between 5 and 30°C temperature, fairing inside air temperature when no active temperature control is and humidity will be measured and re- provided. These ranges will be updated on corded by the ground measurement sys- the basis of thermal analysis results. The tem during the upper composite transporta- thermal analysis will be performed by tion en route to the launch pad and during means of the spacecraft thermal and geo- processing on the launch pad (Table 5-13).

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Mobile thermal Stationary thermal SC batteries system Parameter conditioning unit conditioning unit (optional service)

10..25 10..25 7…20 Temperature of supplied air, °C (adjustable) (adjustable) (adjustable) Dew point of supplied air, °C ≤ -20 Relative humidity of supplied 30* - 60 ≤ 30 air, % Air flow rate, m3/h ≥ 4,000 ≥ 4,000 200 Static overpressure inside 2,000 2,000 12,000 payload fairing, Pa ISO Class 8 ISO Class 8 Cleanliness of supplied air ISO Class 8 (ISO Class 7 optional) (ISO Class 7 optional) * Optionally, a relative humidity of less than 30% can be maintained.

Table 5-12 Air conditioning systems performance data.

Number of Measurement Measurement Parameter sensors range accuracy

Fairing internal air temperature within approx. 2 - 40 … 80°C ± 0.7°C 3 m of fairing separation plane Fairing internal air humidity within approx. 3 m 2 5 … 90 % ± 3 % of fairing separation plane Adapter hardware temperature at 1/2 adapter 2 - 40… 80°C ± 0.7°C height

Table 5-13 Ranges of measured parameters under the payload fairing.

Air conditioning of the upper composite is the required thermal status of the upper provided up to approximately 30 seconds composite is maintained due to the thermal before lift-off and is restarted in about 1 inertia of the thermal cover, the fairing and minute in case of a launch abort. The total the upper composite itself as well as due to number of thermal conditioning interrup- the heat resistance of both the thermal tions is four: cover and the fairing internal thermal blanket. This will be verified by thermal analy- • Upper composite lifting into the service ses carried out for each phase of ground tower processing. • Upper composite installation on the booster unit 5.2.3 In-flight Temperature under the Fairing • Stiffness ring removal The payload fairing protects the payload • Installation of LTC extension during the ascent to a nominal altitude of about 120 km. The duration of each interruption period does not exceed 1 hour. In these periods

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The in-flight adapter hardware temperature flux, terrestrial albedo and terrestrial infra- lies within the range -50°C to +40°C range red radiation. The nominal orientation of without considering the payload-induced the upper composite during the coasting environment. During the ascent, the net phase is described in chapter 3.4. The heat flux density radiated by the fairing does not impact to the spacecraft during the coast- exceed 500 W/m2 at any point. ing phase is determined during the dedicated mission analysis. Within the re- Since the upper stage employs an internal quirements of the upper stage, the mission temperature control system to keep the profile takes into account the spacecraft temperature of the equipment bay that thermal requirements. directly interfaces the payload below 50°C, no induced heat load from here is to be 5.3 Fairing Static Pressure during expected. After the fairing has been jetti- the Ascent soned, the payload is exposed to the Free Molecular Heating (FMH) flux (section The payload compartment is vented during 5.2.5), solar radiation and terrestrial infra- boosted flight. The payload compartment red. pressure and the depressurisation rate are a function of fairing design and flight trajec- 5.2.4 Aerothermal Flux at Fairing tory. The nominal predicted pressure Jettisoning decrease profile for the Rockot payload The time for jettisoning the fairing is deter- fairing is shown in Figure 5-14. The maxi- mined to ensure that the aero-thermal flux mum depressurisation rate will not exceed of 1135 W/m2 will not be exceeded. This 4 kPa/s. The static pressure within the flux is calculated as free molecular heating fairing is measured during flight and trans- acting on a plane surface perpendicular to mitted to the ground using the on-board the velocity direction. telemetry system.

5.2.5 Heat Impact during Coasting Phase

After the fairing has been jettisoned, the spacecraft is exposed to solar radiation

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x 105 1.0 0.9 0.8 0.7 0.6 InnerPressure pressure under under payload fairing fairing 0.5 0.44

Pressure, Pa Pressure, 0.3 Atmospheric pressure 0.2 during flight 0.1 0 0 10 20 30 40 50 60 70 80 90 100 Time, s

Figure 5-14 Variation of fairing inner static pressure during ascent.

5.4 Contamination and Cleanliness 5.5 Electromagnetic Environment

The pyrotechnic systems used for stage, 5.5.1 Launch Vehicle Electro-Magnetic fairing and payload separation are leak Emissions proof and do not cause any organic con- The launch vehicle is equipped with the tamination or debris. Several pyrotechnic following transmission and reception sys- device operations occur during the tems: Rockot/Breeze-KM flight. In all cases, contamination of the spacecraft is avoided • Two telemetry systems with transmit- either by hermetically closed containment ters and antennas, one in the inter- of the pyro charge or via the geometry of stage and one in the second stage the plume relative to the payload. For • A telemetry system with two transmit- payloads that are sensitive to organic ters and antennas in the Breeze-KM stage contamination, a dedicated contamination • A transponder tracking system with a analysis will be performed. The implemen- transmit-receive antenna. tation of the required measures is to be negotiated and will be defined in a con- During on-ground testing and during flight, tamination control plan. All non-metallic the transmission systems create electric materials in the upper composite are se- and magnetic fields. Their characteristics lected according to the Russian GOST are presented in Table 5-14 and Figure standard, which specifies the use of mate- 5-15 that define also the necessary rials with acceptable outgassing properties. immunity of the spacecraft against the ISO Standard 14644-1 Class 8 air cleanli- launch vehicle and cosmodrome emissions ness is provided and continuously moni- during all operations. In all frequency por- tored in the payload preparation rooms and tions not explicitly described in Table 5-14 inside the fairing until lift-off. the electrical field intensity is 90 dBμV/m.

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Max. antenna Calculated level of electrical field Radio Emission emissive power, intensity in adapter plane, dBμV/m transmitter frequency, MHz dBWt with fairing without fairing

Telemetry 1 120 - 130 12.3 107 119 Telemetry 2 1030 - 1050 10.0 105 117 Telemetry 3 1015 -1025 7.8 100 112 Telemetry 4 1015 - 1025 7.8 100 112 20.0 Tracking 2700 - 3000 107 119 (pulsed mode)

Table 5-14 Launch vehicle and cosmodrome electromagnetic emissions.

160 1015 - 1050 MHz 140 120 - 130 MHz 117 dBμV/m 119 dBμV/m 2700 - 3000 MHz 120 119 dBμV/m V/m μ 100 90 dBμV/m

40 Field strength, dB Field strength, 20

0 100 1000 10000 100000 Frequency, MHz

Figure 5-15 Field strength of launch vehicle and cosmodrome electromagnetic emissions in the adapter plane without payload fairing.

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To ensure electromagnetic compatibility 5.5.2 Spacecraft Electro-Magnetic Emissions (EMC) between the launch vehicle and the spacecraft, a frequency plan is prepared In order to avoid electromagnetic interfer- for each launch. The Customer shall ence with the launch vehicle, the space- supply all data needed to support appro- craft radio frequency emission during all priate EMC analyses. operations should not exceed the levels defined in Table 5-15 and Figure 5-16.

Frequency band, MHz Allowable field strength, dBμV/m

120 - 130 80

1015 - 1050 80

2700 - 2900 70

Table 5-15 Allowable spacecraft emissions at Plesetsk Cosmodrome.

160

140 125 dBμV/m 120 V/m μ 100

60 120 - 130 MHz 80 dBμV/m 2700 - 2900 MHz 1015 - 1050 MHz 70 dBμV/m 40 80 dBμV/m Field strength, dB Field strength, 20

0 100 1000 10000

Frequency, MHz

Figure 5-16 Allowable spacecraft emissions at Plesetsk Cosmodrome.

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Chapter 6 Spacecraft Design and Verification Requirements

Table of Contents

6 Spacecraft Design and Verification Requirements...... 6-1 6.1 Safety Requirements...... 6-1 6.1.1 Selection of Payload Materials ...... 6-1 6.2 Design Characteristics ...... 6-1 6.2.1 Mass Properties...... 6-1 6.2.2 Centre of Mass Constraints ...... 6-1 6.2.3 Structural Integrity...... 6-1 6.2.3.1 Factors of Safety...... 6-1 6.2.3.2 Dimensioning Loads ...... 6-2 6.2.4 Stiffness...... 6-2 6.2.5 Overflux ...... 6-3 6.3 Spacecraft Mechanical Qualification and Acceptance Tests...... 6-3 6.3.1 Static Load Test...... 6-3 6.3.2 Sinusoidal Vibration Test ...... 6-3 6.3.3 Random Vibration Test ...... 6-3 6.3.4 Acoustic Noise Test ...... 6-4 6.3.5 Shock Test...... 6-4 6.4 Interface Tests...... 6-4

List of Figures

Figure 6-1 Typical limit load factors for initial dimensioning of secondary structures and equipment brackets...... 6-2

List of Tables

Table 6-1 Sinusoidal vibration loads...... 6-3 Table 6-2 Typical mechanical verification test matrix for qualification (Q) and acceptance (A) of the spacecraft...... 6-4 Table 6-3 Acoustic noise spectrum...... 6-4

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6 Spacecraft Design mass moment of inertia better than ±10 %. The CoM shall be specified with an accu- and Verification racy of 50 mm along the launch vehicle Requirements longitudinal axis and within a circle of 30 mm radius around the launch vehicle longitudinal axis. This chapter defines the spacecraft design and verification requirements that have to For spacecraft using liquid propellant the be taken into account in preparation of a dynamics of the liquid shall be specified by launch on Rockot/Breeze-KM. Any devia- means of a proper sloshing model at dif- tion from these requirements has to be ferent acceleration levels. mutually agreed. 6.2.2 Centre of Mass Constraints 6.1 Safety Requirements The Rockot launch vehicle is capable of The Customer is required to design and supporting a large variation of the CoM operate the spacecraft in accordance with position along its X-axis. However, the the launch site safety regulations described dependency of the lateral accelerations on in chapter 9. It must be assured by appro- the CoM position may limit its position on priate means (mechanical ground support the X-axis. equipment design, operational procedures) The total displacement of the composite that constraints related to ground opera- CoM of the spacecraft or a combination of tions do not become design drivers for the multiple spacecraft and the Breeze-KM flight hardware. must stay within a radius of 30 mm around the launch vehicle longitudinal axis. This 6.1.1 Selection of Payload Materials imbalance directly affects the controllability of the upper stage and thus the spacecraft Properties as well as types of materials angular velocities at separation and components used for the spacecraft design must be based on recognised stan- 6.2.3 Structural Integrity dards agreed by the launcher authority.

6.2.3.1 Factors of Safety 6.2 Design Characteristics The minimum factors of safety to be taken 6.2.1 Mass Properties into account for structural dimensioning are: The spacecraft mass properties shall be defined according to the following accura- • Yield load: j ≥ 1.1 cies to enable dynamic analyses to be • Ultimate load: j ≥ 1.25 undertaken as part of the overall spacecraft to commence vehicle preliminary mis- The factors of safety apply to combinations sion analyses. The mass shall be specified of simultaneously acting mechanical and with an accuracy of better than ±2.5 %, the thermal limit loads.

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1000

100 Limit load factor, g loadLimit factor,

10 0.1 1 10 100

Equipment mass, kg

Figure 6-1 Typical limit load factors for initial dimensioning of secondary structures and equipment brackets.

6.2.3.2 Dimensioning Loads Limit load factors cover equipment / unit responses due to quasi-static and low- Structural dimensioning must take into frequency transient and random accelera- account critical combinations of simultane- tions encountered during lift-off and ascent. ously acting load types.

Generic design accelerations for space- 6.2.4 Stiffness craft primary structure dimensioning are compiled in chapter 5.1.2. To avoid dynamic coupling between the low-frequency launch vehicle and payload Secondary structures and equipment modes, the payload minimum natural fre- brackets must be dimensioned taking into quency f0 is required to be: account local responses to the combined effect of simultaneously acting low fre- • Lateral (Y-Z plane): f0 ≥ 15 Hz quency transient and high frequency ran- • Axial (X): f0 ≥ 33 Hz dom vibrations. Typical mass-dependent combined load factors n are presented in Note: Resonance requirements are related Figure 6-1 as a design guideline. For di- to spacecraft modes with significant effec- mensioning, limit load factors have to be tive mass of meff ≥ 70%. The minimum applied natural frequency values are targets for design, if existing spacecraft are not com- • at equipment / unit CoM, pliant, the given requirement can be re- • in the worst case spatial direction with laxed based on CLA results. For heavy respect to resulting stresses and reac- spacecraft, additional coordination with tions. EUROCKOT is required to ensure that the structural modes of the integrated space-

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craft/ adapter stack do not decrease below • Qualification model: ultimate load (1.25 critical limits. times limit loads)

• Protoflight model: yield load (1.1 times 6.2.5 Overflux limit load) Overflux refers to disturbances of the axial For a realistic simulation of the load intro- line load at the interface of the adjacent duction, the spacecraft must be attached to mating structures. These local distur- a flight representative adapter or separa- bances are caused by structural disconti- tion system during the static test. nuities such as stringers, cut-outs, etc.

Overflux requirements apply to clamp 6.3.2 Sinusoidal Vibration Test adapters only and will be specified on a The inputs at the spacecraft adapter inter- case-by-case basis. face are shown in Table 5-2, and the test factors in Table 6-1. 6.3 Spacecraft Mechanical Qualification and Acceptance A permission for notching of sine test lev- Tests els at critical resonances may be requested from EUROCKOT in order not to The Customer shall demonstrate that the exceed the spacecraft flight responses spacecraft structure complies with the re- predicted by coupled load analysis. quired design characteristics as defined in Chapter 6.3, taking into account the envi- Acceptance Qualification ronmental conditions stated in chapter 5. Test factors 1 1.25 Additionally, spacecraft mathematical mod- Sweep rates (one sweep per 4 oct/min 2 oct/min els submitted to the launcher authority for axis) performance of final coupled analyses and Table 6-1 Sinusoidal vibration loads. flight mechanics analyses must be verified by tests. 6.3.3 Random Vibration Test A typical qualification/acceptance test matrix is shown in Table 6-2. The spacecraft Random vibration test is recommended verification plan finally selected needs to only for small satellites of 100 kg mass or be approved by the launcher authority. less and for satellites with small dimensions. The vibration loads for this purpose 6.3.1 Static Load Test will be specified on a case-by-case basis, see chapter 5.1.5. On the basis of dimensioning loads given in chapter 6.2.3.2, EUROCKOT defines For larger spacecraft, EUROCKOT rec- critical load cases to which the spacecraft ommends to perform an acoustic test to structure will be subjected. The structure accurately reflect the in-flight random envi- must successfully pass static load tests up ronments experienced. Since the vibration to: level depends on the dynamic properties of

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the payload adapter structure, this test A permission for notching of random vibra- should be performed with the spacecraft tion test levels at critical resonances may attached to a flight-like payload adapter be requested from the launcher authority in (not hard mounted) to accurately represent order not to exceed local responses meas- the flight configuration. ured during an acoustic noise test or determined by an acoustic response analysis.

Required Tests

Static Sinusoidal Random Acoustic Shock Test Hardware Chap. 6.3.1 Chap. 6.3.2 Chap. 6.3.3 Chap. 6.3.4 Chap. 6.3.5

Q A Q A Q A Q A Q A

Prototype Philosophy: Qualification Model X X X X1) X1) X X X X2) Flight Model

Protoflight Philosophy: X X X1) X X2) Protoflight Model 1) alternatively for small satellites 2) optionally Table 6-2 Typical mechanical verification test matrix for qualification (Q) and acceptance (A) of the spacecraft.

6.3.4 Acoustic Noise Test 6.3.5 Shock Test

The acoustic noise spectrum as defined in Shock tests of complete spacecraft must chapter 5.1.4 must be used as the test be conducted by firing of the planned sepa- input with the factors and durations of ration system. For predicted shock re- Table 6-3 applied. sponse spectra, see Chapter 5.1.6.

Accep- Qualifi- Protoflight tance cation Qualification 6.4 Interface Tests

Test factor for Per SC designer’s national stan- Depending on the specific mission the fol- acoustic dard pressure lowing set of compatibility tests may have to be performed: Exposure 60 120 60 duration, s • The Matchmate test, also referred to as Table 6-3 Acoustic noise spectrum. Fit-check test for verification of electrical and mechanical interfaces of the The requested test duration takes into ac- spacecraft to the adapter and separa- count a scatter factor significantly greater tion system. This test is performed than four. preferably with flight units and can be combined with a functional test of the

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separation system. This test is manda- tory for each mission.

• If necessary, volume compatibility test with fairing and adapter. A satellite dummy simulating the spacecraft static envelope will be used for this purpose

Additionally, the following test have to be performed in case of necessity:

• Thermal tests

• EMC tests

• Dedicated electrical interface tests when non-standard interfaces are used

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Chapter 7 Mission Management

Table of Contents

7 Mission Management ...... 7-1 7.1 Mission Management Overview ...... 7-1 7.2 Organisation and Responsibilities ...... 7-1 7.2.1 EUROCKOT Mission Responsibilities ...... 7-3 7.2.1.1 Mission Integration...... 7-3 7.2.1.2 Interface Design, Qualification and Verification ...... 7-4 7.2.1.2.1 Design of the Payload Adapter ...... 7-4 7.2.1.2.2 Payload Adapter Qualification Test at KSRC...... 7-4 7.2.1.2.3 Fit Check of FM Spacecraft with FM Payload Adapter ...... 7-5 7.2.1.3 Configuration Control...... 7-5 7.2.1.4 Launch Vehicle Procurement...... 7-5 7.2.1.5 Spacecraft Preparation and Launch Operations...... 7-5 7.2.1.6 Post-Launch Activities...... 7-6 7.2.1.7 Quality and Mission Assurance...... 7-6 7.2.1.8 Safety Provisions ...... 7-6 7.2.1.9 Risk Management...... 7-6 7.2.1.10 Technology Transfer and Security...... 7-7 7.2.2 Customer Mission Responsibilities ...... 7-7 7.3 Reviews and Documentation...... 7-8 7.3.1 EUROCKOT Documents ...... 7-10 7.3.1.1 Interface Control Document...... 7-10 7.3.1.2 Preliminary Design Review Data Package ...... 7-10 7.3.1.3 Critical Design Review Data Package ...... 7-10 7.3.1.4 Joint Operations Plan ...... 7-11 7.3.1.5 Safety Reply ...... 7-11 7.3.1.6 Launch Evaluation Report ...... 7-11 7.3.2 Customer Documents ...... 7-11 7.4 Overall Mission Schedule...... 7-12

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List of Figures

Figure 7-1 Industrial organisation of EUROCKOT and its major subcontractors...... 7-2 Figure 7-2 EUROCKOT mission management organisation...... 7-3

List of Tables

Table 7-1 Typical launch services reviews and meetings...... 7-9 Table 7-2 Typical ICD structure...... 7-10 Table 7-3 Documents to be supplied by the Customer...... 7-12 Table 7-4 Typical mission schedule...... 7-13

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7 Mission 7.2 Organisation and Management Responsibilities EUROCKOT is responsible to the Cus- 7.1 Mission Management tomer for all commercial and technical ac- Overview tivities within the launch contract. EUROCKOT implements this contract as Mission management is conceived by the single prime contractor towards the EUROCKOT to fulfil all the Customer's Customer and as the Customer’s sole in- requirements to the greatest possible ex- dustrial partner for all aspects of the law. tent by undertaking the following activities: EUROCKOT is a company governed by German law and offers all the legal safe- • Management and planning of the entire guards provided by a Western company. mission integration process As a constituent company of EUROCKOT, • Definition and control of all payload/ Khrunichev State Research and Production launch vehicle interfaces Space Center (KSRC) of Russia provides • Performance of mission design and the launch vehicle as well as the launch mission analyses services and launch operations support through a sub-contract to the Russian • Provision of the launch vehicle with Space Forces. appropriate interfaces including the payload adapter and separation system EUROCKOT's other parent company Astrium which is located next door to • Transport of the customer's spacecraft EUROCKOT, offers support in engineering and support equipment from the Rus- and commercial areas as necessary. The sian port-of-entry to EUROCKOT's fa- distribution of the relevant activities among cilities at the launch site EUROCKOT, KSRC and Astrium is de- • Provision of appropriate payload prepa- picted in Figure 7-1. ration facilities within EUROCKOT's fa- For mission management, EUROCKOT cilities at the launch site uses a scheme (Figure 7-2) which has • Management and performance of pre- proven extremely successful in the past. launch operations and launch Customers conclude a Launch Services Agreement (LSA) directly with • Performance of trajectory tracking and EUROCKOT, who provides the single point payload telemetry data reception, as of focus for the Customer through a desig- required nated Mission Manager. The Mission Man- ager is responsible for the management of all the launch service tasks. The Mission Manager has full programme authority and is responsible for all coordination required to implement the launch contract. The Mis-

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sion Manager is responsible for ensuring technical team. The Mission Manager re- that all payload launch requirements are ports to the Technical Director and the met and is in continuous contact with the CEO. Within each individual project, the Customer from contract signature up to Mission Manager is supported by another launch. member of the technical team, thereby ensuring common standards and practices At the launch site, he/she acts as the day- within EUROCKOT as well as providing to-day intermediary between the Customer personnel redundancy. He/she is sup- and the launch site authorities for the pur- ported by other members of the pose of satisfying the Customer’s require- EUROCKOT team to fulfil contractual obli- ments. This includes responsibility for gations including the contracts and finance launch operations planning, procedures team which is responsible for all contrac- and launch execution. The launch decision tual, commercial and financial matters and is the responsibility of a Management the sales team for public relations activi- Group consisting of the EUROCKOT Mis- ties. sion Manager and representatives of KSRC, the Customer and the launch site Within EUROCKOT, the Mission Manager authorities. represents the interests of the Customer, towards the Customer he represents the The Mission Manager acts as part of a interests of EUROCKOT. The structure of team of experienced programme manag- the mission management organisation and ers, mission managers and engineers who its relationship to the Customer and KSRC make up the nucleus of the EUROCKOT is shown in Figure 7-2.

KHRUNICHEV State Research and Production Space Center

Prime contractor

Launch vehicle and services Launch operations at Plesetsk Logistic support procurement Cosmodrome

Technical and hardware Flight evaluation management Safety Insurance Quality Financial arrangements Mission assurance (reliability)

Figure 7-1 Industrial organisation of EUROCKOT and its major subcontractors.

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7.2.1 EUROCKOT Mission annexes of the launch contract. This ICD is Responsibilities agreed and signed by all parties including EUROCKOT, Khrunichev, the Customer EUROCKOT will manage all mission- and the Spacecraft developer. Preliminary related activities from first preliminary esti- mission design and mission analyses are mations before launch contract signature then performed versus the requirements through post-launch evaluation and review contained within the ICD and presented in with the emphasis on Customer satisfac- the Interface Preliminary Design Review tion. (PDR) for customer review and approval. After the PDR the ICD is updated and put 7.2.1.1 Mission Integration under formal configuration control, with EUROCKOT responsible for maintenance EUROCKOT's mission integration respon- and updating of this document. When the sibility includes the definition and control of spacecraft design matures and the final the spacecraft to launch vehicle interfaces spacecraft data is known the final mission as well as the performance of the mission design and mission analyses are then per- design and mission analyses (Chapter 8). formed versus the requirements contained Initially a draft Interface Control Document within the ICD and presented in the (ICD) which contains the requirements and Interface Critical Design Review (CDR) for design solutions for the interfaces is estab- customer review and approval. lished at the start of the mission integration phase. This is based upon customer re- The ICD is maintained under formal con- sponses to the questionnaire "SC Initial figuration control until launch. This docu- Data Requirements for the Rockot launch ment also includes relevant technical vehicle", ESPE-0022, as well as the Inter- specifications relating to payload prepara- face or Technical Requirements Document tion facilities at the range. (IRD) which forms part of the technical

Customer

EUROCKOT Mission Manager

Astrium ST

Contracts and Technical Management/ Sales Team Finance Team Team Technical

Support

Khrunichev State Research and Production Space Center

Figure 7-2 EUROCKOT mission management organisation.

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flight units are produced together. On the 7.2.1.2 Interface Design, Qualification and Verification basis of the existence of a fully qualified launch vehicle including the Breeze-KM The design of the interfaces between the upper stage and payload fairing, only the launch vehicle upper stage and the space- mission-specific interface has to be veri- craft, i. e. the payload adapter, attachment fied. Depending on the degree of individu- and separation system, is one of the first ality of the payload adapter requested to activities to be started after technical kick- fulfil specific spacecraft designs, a qualifi- off. In the majority of cases, a mission- cation test program will be set up. specific payload adapter design is used, due to the fact that even for similar designs 7.2.1.2.2 Payload Adapter Qualification Test at there can still be small but significant dif- KSRC ferences between designs, e. g. connector The qualification programme of the pay- type and layout, spring pusher location and load adapter will typically consist of some forces, clamp band pre-tension, adapter of the following test steps: geometry and height as well as different spacecraft mass properties etc. Hence, in • Static tests the cases where qualification by similarity is not applicable, a design and qualification • Dynamic tests process will be undertaken to cover these • Spacecraft separation tests differences. • Fit-check with payload adapter or op- 7.2.1.2.1 Design of the Payload Adapter tionally, in case of low clearances to the payload fairing, a volume fit-check Compliance of the design with the Cus- with the fairing tomer requirements stated in the ICD and with the environmental constraints is dem- The tests are performed at the premises of onstrated in the PDR and CDR. A success- KSRC in Moscow, Russia. The tests are ful CDR signifies the payload adapter de- performed using a Breeze-KM upper stage sign approval and the go-ahead for manu- mechanical interface simulator together facturing. If, on the other hand, design with a spacecraft Mass Frequency Simula- changes are necessary because of the tor model (MFS). The MFS is a high fidelity CDR, this go-ahead is given when these reproduction of the spacecraft flight model design changes are successfully com- electro-mechanical interfaces and also pleted. reproduces the major mass and dynamic properties of the spacecraft including its However, owing to time constraints, the lowest natural frequencies. In case of low procurement of long lead items and initia- clearances between the spacecraft and tion of piece part manufacturing can start payload fairing, an additional volume fit earlier. Like the necessary effort for qualifi- check, using a spacecraft volume simulator cation, this depends mainly on the degree can also be undertaken as an option. of individuality of the payload adapter for each mission. Ideally, qualification and

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The necessity and extent of each test de- 7.2.1.3 Configuration Control pend mainly on spacecraft mass and stiff- Programme-specific configuration control ness properties, and geometrical interfaces and data management procedures begin environment sensitivities. immediately upon contract signature with To further verify correct attachment inter- EUROCKOT and cover all documents and face and integration feasibility, two addi- data exchanged with the Customer. These tional means to confirm interface compati- documents are defined in chapter 7.3. The bility can be undertaken and are described overall programme configuration control of in the following paragraphs. the launch vehicle and launch service is an extension of the KSRC Quality and Mission 7.2.1.2.3 Fit Check of FM Spacecraft with FM Assurance Plan. Data management is an Payload Adapter integral part of this plan. A unique programme configuration control plan is pre- A fit check, preferably using the flight pared. This plan shows: model spacecraft and the flight model payload adapter or an identical model thereof, • Responsibilities for configuration man- is performed at the spacecraft manufac- agement in each organisation turer’s premises demonstrating that the mating and separation of mechanical and • Documentation subject to configuration electrical joints comply with the ICD re- management quirements. Depending on the customer's • Change orders issued specific requirements, a separation test involving ignition of the separation system • Orders processed pyrotechnics can be undertaken to verify • Constitution of the joint change board shock levels at the spacecraft interface.

7.2.1.4 Launch Vehicle Procurement 7.2.1.2.4 Master Gauge Interface Verification EUROCKOT monitors the launch vehicle A master gauge / drill template, produced production progress according to the mis- by either the spacecraft contractor or sion integration schedule, approves launch EUROCKOT, ensures that the correct posi- vehicle acceptance and is responsible for tions of the fixing points are achieved not the specification and provision of the only by compliance with the interface draw- launch vehicle to spacecraft interfaces and, ings but also by using the identical tool for if applicable, adaptation of the fairing. applying them on the test and flight hardware. This is generally used for point at- 7.2.1.5 Spacecraft Preparation and Launch tachment systems to ensure correct posi- Operations tioning of the interface points on the spacecraft for the separation system. EUROCKOT defines the launch operations for the launch site through the establishment of a Joint Operations Plan (JOP) for all joint activities with the customer. This includes all activities which involve support

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from EUROCKOT and KSRC at the launch A system of procedures which has proven site, for example launch site support activi- its efficiency in the past assures detailed ties including fuelling support as well as analysis of discrepancies as well as related combined operations of the launch vehicle dispositions and verification of their execu- and spacecraft. The plan defines the tion. launch site organisation, joint operations, facilities, operational support, electrical 7.2.1.8 Safety Provisions check-out as well as launch day activities including go/no-go criteria. The Spacecraft EUROCKOT will provide the single focal Operation Plan (SOP) (section 7.2.2) pro- point for all system and range safety mat- vided by the Customer provides the inputs ters. The Customer will provide the neces- for this plan and will be adhered to. The sary data as described in chapters 9 and JOP is the working document used by 12 to enable EUROCKOT to obtain space- EUROCKOT, the Customer, and KSRC craft safety approval for the launch cam- and its subcontractors to manage the paign. launch campaign. 7.2.1.9 Risk Management

7.2.1.6 Post-Launch Activities Risk management by EUROCKOT covers State vectors of the upper stage at burn- the following risk in particular: out and of the satellite separation event will Political Risk: The EUROCKOT pro- be provided as preliminary data approxi- gramme is part of the German-Russian mately 60 minutes after separation. Two space cooperation agreement, backed by months after launch, EUROCKOT provides high level guarantees from the German a Launch Evaluation Report (LER), show- and Russian Governments, explicitly in- ing the performance achieved and the be- cluding the Russian Space Agency and the haviour of the launch vehicle. This report is Space Forces. based on processed launch vehicle telemetry and tracking data as well as on Commercial Risk: Astrium is financially spacecraft orbit data provided by the Cus- backing all required funding. tomer. Technical Risk: The Rockot launch vehicle is fully operational. The first two stages 7.2.1.7 Quality and Mission Assurance undergo extensive tests (DPA / test firing) At KSRC, the Deputy Director for Quality on a yearly basis. The commercial Rockot Assurance, reporting directly to the Gen- configuration, including Breeze-KM and the eral Director, ensures and supervises large payload fairing, was successfully compliance with all relevant requirements. flight-qualified for the first time in May Quality audits and procedures maintain 2000. The co-production of Breeze-KM for rigorous adherence to all elements in the Rockot and Breeze-M for Proton ensures factory and launch site operations. Incom- programme continuity at KSRC. ing materials and subcontractors are certi- Launch Risk: Launchers which are held in fied, continuously reviewed and inspected. stock for rapid replenishment ensure short

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reaction launches in the case of satellite fense Technology Security Administration problems and immediate re-launch in the (DTSA). case of launch failure after completion of a For spacecraft coming under DTSA juris- failure investigation. In the unlikely event of diction, special measures will be taken to a launch failure, a contingency plan previ- meet these requirements. In the case of ously reviewed and tailored to the specific technology transfer issues, it is recom- mission would control the total process mended that a Technical Assistance providing a failure action list from data re- Agreement (TAA) with DTSA is concluded view through anomaly identification and the very early on in the programme to allow for setting up of analysis and review boards up technical interchanges between the space- to final explanation and corrective action craft contractor and the launch service dispositions. companies, e.g. EUROCKOT, KSRC and Almost 1500 launches from Plesetsk com- their subcontractors. Physical security of bined with the Proton experience of KSRC the spacecraft, of its associated support and the Ariane experience of Astrium fur- hardware and of documentation at the ther reduce technical risks for the Cus- launch site is assured by physical barriers tomer. such as controlled entry doors, round-the- clock surveillance of the hardware by secu- 7.2.1.10 Technology Transfer and Security rity guards and agreed procedures.

EUROCKOT is committed to meeting gov- 7.2.2 Customer Mission Responsibilities ernment and Customer requirements concerning technology transfer issues and the The Customer is required to designate a physical security of the spacecraft, its sup- Payload Mission Manager who will be the port equipment and associated documen- single point of contact for the Mission Man- tation during the mission integration proc- ager at EUROCKOT. Early in the contract ess and the launch campaign. For this implementation process, the Customer is purpose, the mission integration process requested to provide responses to the and launch site activities conducted by questionnaire "SC Initial Data Require- EUROCKOT, for instance all technical ments for the Rockot Launch Vehicle", interchange meetings (TIM), data transfer ESPE0022" which covers: from the spacecraft contractor, e. g. drawings and mathematical models, and activi- • Required mission characteristics ties at the launch site will be governed by a • Spacecraft characteristics (dimen- specific document generated for PDR and sional, electrical, thermal, environ- updated for CDR. mental, etc.)

For the majority of spacecraft contractors, • Spacecraft launch preparations re- this document will be based to a large ex- quirements tent on United States requirements issued by the Office of the Secretary of the De- Early in the mission analysis process, the Customer submits a payload development and test plan to meet the Rockot/Breeze-

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KM environmental conditions. Additionally, 7.3 Reviews and the Customer shall provide several space- Documentation craft software models, especially for integrated structural and thermal analyses Within each phase of the launch service (chapter 8 and section 12.2). During the implementation there are various activities mission design and mission analysis proc- and milestones planned to enable success- ess, the Customer is requested to submit ful fulfilment of the contract. These activi- environment test results (section 12.4). ties include regular meetings with the cus- The Customer attends the preliminary and tomer and spacecraft contractor and also final mission analysis reviews which are the generation of documents and analyses undertaken as part of the Launch Vehicle for review and approval. The activities are Preliminary and Critical Design Review. coordinated by the EUROCKOT Mission Manager at the start of the contract. As an input to the setup of the Joint Opera- tions Plan (JOP), the Customer will issue Typically EUROCKOT tries to distribute the Spacecraft Operations Plan (SOP). For meetings approximately evenly between the operational activities at the range, the the customer, spacecraft contractor and Customer will provide procedures for the EUROCKOT / KSRC sites. Generally, to various operations on the spacecraft for have easy access to specialists, the main safety examination by the range authority mission design and mission analysis re- (see Section 12.5). views, the Preliminary Design Review (PDR) and the Critical Design Review A safety review based on three safety sub- (CDR) are held at KSRC's premises in missions (phases I, II and III) and the Moscow. Other technical interchange spacecraft safety certificate provided by meetings and reviews can be held at other the Customer or spacecraft contractor locations including the customer site de- must also be completed during the launch pending on the specific purpose. A sum- preparation phase. A phase I safety sub- mary of reviews and their typical allocation mission is expected at the start of the con- within the mission schedule is given in tract phase. For details, see chapter 9 and Table 7-1. section 12.3. The aim is, where possible, to combine Hardware models which have to be pro- some of these meetings and reviews in vided, especially a spacecraft mass fre- order to optimise the time and cost in- quency simulator model and, if necessary, volved for all parties. An overall summary the volume fit-check dummy are described of documents to be supplied by in sections 7.2.1.2 and 12.7. In general, all EUROCKOT and the Customer, as well as items to be provided by the Customer such their typical release dates is given in sec- as documents, software and hardware tion 7.4. models are summarised in chapter 12.

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Meetings / reviews schedule Date

Contract signature meeting L - 24 months Technical kick-off meeting/ IRD review L - 24 months Launch vehicle to spacecraft system requirements review + ICD outline L - 22 months ICD review (draft issue) L - 21 months Launch Vehicle to Spacecraft Preliminary Design Review incorporating the L - 18 months preliminary mission design and analyses ICD review (issue 1) L- 12 months Spacecraft Operations Plan/ Joint Operations Plan review As necessary, combined with other meetings Technical interchange meetings As necessary, combined with other meetings Safety reviews (phases I, II and III) As necessary, combined with other meetings Launch Vehicle to Spacecraft Critical Design Review incorporating the final L - 12 months mission design and analyses ICD review (issue 2) L - 11 months Design qualification review L - 5 months ICD review (final issue, if applicable) L - 5 months Shipment Readiness Review To be agreed Launch Readiness Review (LRR)/ State Commission L - 3 days Launch Evaluation Review meeting L + 2 months

Table 7-1 Typical launch services reviews and meetings.

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7.3.1 EUROCKOT Documents 7.3.1.2 Preliminary Design Review Data Package The main documents to be established by Preliminary Design Review Data Package L-19 months EUROCKOT are summarised below: The results of the preliminary mission 7.3.1.1 Interface Control Document design and the analyses are documented in the preliminary design review data pack- Interface Control Document L - 21 to 5 months age. This includes: The ICD is the document that guarantees • Payload accommodation design includ- the technical definition and control of all ing mission-specific equipment and in- interfaces between the launch system and terfaces the payload to both the spacecraft Cus- tomer and to EUROCKOT. In addition, the • Trajectory and mission sequence ICD is intended to establish the operational • Spacecraft separation analysis requirements for a launch campaign. The • Ground and flight thermal environment document will be updated regularly with • Dynamic Coupled Loads Analysis inputs from the Customer and updated by • Cleanliness the EUROCKOT Mission Manager in agreement with all parties. The ICD is a • Measurement system living document, being constantly updated • Telemetry to reflect the latest status of the launch • Radio frequency compatibility services. A typical ICD structure is de- • Electrical picted in Table 7-2. • Pre-launch support and operations 1 Introduction • Reliability 2 Mission Requirements • Social services 3 Mechanical and Electrical Interfaces • Communication infrastructure 4 Mechanical Loads and Environments • Security 4.1 Flight Loads • Transportation 4.2 Ground Loads

4.3 Thermal 7.3.1.3 Critical Design Review Data Package 4.4 Cleanliness Critical Design Review Data Package L - 13 months 4.5 EMC The results of the final mission design and 5 Preparation Facilities analyses are documented in the critical 5.1 General Requirements design review data package. This includes: 5.2 Communication • Payload accommodation design includ- 6 Verification Matrix ing mission-specific equipment and in- Table 7-2 Typical ICD structure. terfaces • Trajectory and mission sequence • Spacecraft separation analysis

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• Ground and flight thermal environment 7.3.1.5 Safety Reply • Dynamic coupled loads analysis / loads Safety Reply, Phases I; II; III L – 17; 11; 4 months • Cleanliness • Measurement system The EUROCKOT safety reply contains: • Telemetry • Assessment of customer safety sub- • Radio frequency compatibility missions • Electrical • Description of spacecraft systems and classification of hazardous systems • Pre-launch support and operations • Development of list of hazards and • Reliability reports on analysis of these hazards • Social services • Verification of spacecraft design com- • Communication infrastructure pliance with the standards of either the • Security country of origin or agency • Transportation • Safety constraints tailored to dedicated spacecraft

7.3.1.4 Joint Operations Plan • Verification of spacecraft design com- Joint Operations Plan L - 13 months pliance with EUROCKOT Safety Hand- book EHB0004 and establishment of The JOP covers all joint operations involv- reports on non-compliance with these ing EUROCKOT, Khrunichev including provisions subcontractors and the Customer, such as • Specific verification guidelines are gi- joint activities where launch site support is ven in chapter 9 needed like spacecraft fuelling as well as combined operations of the launch vehicle 7.3.1.6 Launch Evaluation Report and spacecraft from the beginning of encapsulation to lift-off. The JOP also in- Launch Evaluation Report L + 2 months cludes the agreed go/ no-go criteria for The launch evaluation report provides: launch. • Launch vehicle performance based on Note: Specific spacecraft operations are telemetry and tracking data the responsibility of the Customer and to be included in the SOP. • Launch vehicle behaviour • Launch vehicle / spacecraft interface aspects including launch environment

7.3.2 Customer Documents

The documents to be provided by the Cus- tomer are described in detail in chapter 12 and summarised in Table 7-3.

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Date (typically) Documents to be provided Preliminary Final

Interface Requirements Document (IRD) L - 24 months I: L-18 months Safety Submission (Phase I, II, III) III: L- 6 months II: L-12 months SC Safety Certificate L – 12 months III: L – 5 months SC Flight Readiness Certificate L – 6 months L – 5 months Flight Readiness Data Package L – 6 months L – 5 months Spacecraft Mechanical Environment Test Plan L - 16 months Spacecraft Dynamic Model (Preliminary) L - 23 months L - 11 months Spacecraft Thermal Model (Preliminary) L - 23 months L - 11 months Response to Questionnaire: Input to Mission Design and Mission Analysis L - 23 months L - 11 months Spacecraft Operations Plan L - 17 months Spacecraft Mechanical Environment Qualification Test Results L - 6 months Final Spacecraft Mass Properties L - 7 days Orbital Tracking Operation Report L + 2 weeks

Table 7-3 Documents to be supplied by the Customer.

Table 7-4. Nevertheless, the mission- 7.4 Overall Mission Schedule specific schedule will be established during The overall mission planning is designed to the mission kick-off meeting. It will address provide the Customer with a reasonably in particular the spacecraft development minimum lead time of 24 months from con- and qualification schedule as well as other tract signature to launch, while still allowing Customer requirements. If the results of for thorough technical preparation, in par- the preliminary mission analysis are re- ticular through mission analysis. If a repeat quired for spacecraft development and launch for a similar spacecraft and compa- qualification earlier than indicated in this rable orbit characteristics is requested by section, a project begin earlier than L-24 the Customer, lead times of 15 months months can easily be implemented. The should be achievable and would be the schedule of the launch campaign is de- subject of specific agreements. scribed in chapter 10.

The mission schedule of a typical mission with a 24 month lead time is depicted in

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Table 7-4 Typical mission schedule.

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Chapter 8 Mission Design and Mission Analysis

Table of Contents

8 Mission Design and Mission Analysis...... 8-1 8.1 General...... 8-1 8.2 Overall Description of Spacecraft to Rockot Launch Vehicle Integration (Book 1)...... 8-3 8.3 Trajectory and Mission Sequence (Book 4) ...... 8-3 8.4 Dynamics of Spacecraft Separation (Book 5) ...... 8-3 8.5 Thermal Environment (Book 6, parts 1 and 2) ...... 8-4 8.6 Dynamic Coupled Loads Analysis (Book 7) ...... 8-5 8.7 Spacecraft Cleanliness Control (Book 8) ...... 8-6 8.8 Measurement System (Book 9)...... 8-7 8.9 Electromagnetic Compatibility Study (Book 10) ...... 8-7 8.10 Onboard Electrical Interface of Spacecraft and Launch Vehicle (Book 11) ...... 8-8 8.11 Ground Electrical Cabling, Power Supply and Interfaces with Spacecraft GSE (Book 12) ...... 8-8 8.12 Launch Base Operations Support (Book 14 part 1, 2, 3) ...... 8-8 8.13 Reliability (Book 15, part 1) ...... 8-9 8.14 ILV Components Quality Assurance (Book 15, part 2)...... 8-9 8.15 Social services (Book 16)...... 8-9 8.16 Communications support (Book 17) ...... 8-9 8.17 Security (Book 18)...... 8-9 8.18 Transportation from Port-of-Entry to the Launch Site Facilities (Book 19) ...... 8-10

List of Tables

Table 8-1 Data package structure...... 8-2 Table 8-2 Data package book content...... 8-2

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8 Mission Design and and to an update in the input data from the customer. Mission Analysis • Final mission design and analysis. This 8.1 General uses final input data from the customer to finalise and freeze the actual design Consistent with the dedication to provide for the launch campaign and flight. The the highest standards and a successful data package is reviewed in the launch launch of the customer spacecraft, vehicle to spacecraft Critical Design EUROCKOT strives to conduct a thorough, Review (CDR). Following successful detailed and transparent mission design conclusion of the CDR the design for and analysis process for the Customer. flight is formally released. This is evidenced by the extensive and detailed data packages provided to the The contents and tasks undertaken for the customer during the review process, as preliminary and final mission design and described later in this chapter. analyses are identical and differ only in the updated input data. The results of the Within the framework of mission integration design and analyses are presented in the activities, mission design and mission form of books. An overview of the data analyses are conducted to ensure that the package structure is provided in Table 8-1 customer's mission objectives can be below. achieved, e. g. reliable spacecraft injection into the required orbit in the correct atti- Each book contains the structure as shown tude, provision of facilities meeting the in Table 8-2. The contents of the individual Customer's requirements,. The design and books are summarised in the next sec- analyses are conducted on the basis of tions. inputs from the Customer (chapter 12) and the compatibility checked versus the Inter- face Control Document requirements. They are undertaken in two phases:

• Preliminary mission design and analy-

sis. This uses preliminary input data from the customer to confirm the basic design and mission scenarios to the customer. The data package is reviewed in the launch vehicle to space-

craft Preliminary Design Review (PDR). Following a successful PDR, the input data and mission aspects are studied and refined, leading to an update of the Interface Control Document

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Book Content

Book 1 Overall Description of Rockot Launch Vehicle Integration to Spacecraft Book 2 Design Data Package Structure - for internal KSRC use only. Not provided to the customer. Book 3 Input Data (ICD based) - for internal KSRC use only. Not provided to the customer. Book 4 Trajectory and Mission Sequence Book 5 Spacecraft Separation Dynamics Book 6, part 1 Thermal Analysis: Ground Processing Book 6, part 2 Thermal Analysis: Flight Book 7 Coupled Loads Analysis Book 8 Spacecraft Cleanliness Control Book 9, part 1 Measurement System Book 9, part 2 Launch Information Telemetry and Navigation Ballistic Support Book 10 Spacecraft Electromagnetic Compatibility with the Launch Vehicle and Launch Site Book 11 Onboard Electrical Interface of Spacecraft and Launch Vehicle Book 12 Ground Electrical Cabling, Power Supply and Interfaces with Spacecraft GSE Book 13 Intentionally left blank Book 14, part 1 Launch Base Operations: Operations Flow Book 14, part 2 Launch Base Operations: Facilities and GSE (except fuelling support) Book 14, part 3 Launch Base Operations: SC Fuelling Support Activities and Facilities Book 15, part 1 Reliability Book 15, part 2 ILV Components Quality Assurance Book 16 Social Services (Hotel, Intra launch site transportation etc) Book 17 Communications Support Book 18 Security Book 19 Transportation of SC and Ground Support Equipment to Launch Site Facilities

Table 8-1 Data package structure.

Section Titles Contents

Requirements Verification • Per ICD and contract requirements Matrix • Summarized results of analyses and design work Results • Demonstration that ICD requirements are met • Outstanding problems Conclusions • Requirements remaining to be met • Actions to meet requirements • Description of techniques used for the analyses Detailed Data • Input data • Miscellaneous

Table 8-2 Data package book content.

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• launch window constraints from the 8.2 Overall Description of spacecraft side. Spacecraft to Rockot Launch Vehicle Integration (Book 1) The trajectory analysis provides a description of the launch vehicle during stage 1, This book provides a top level description stage 2 and upper stage flight phases. This of the measures taken to integrate the includes velocity, altitude, dynamic pres- customer's satellite to the Rockot/Breeze- sure, flight angle and position of the KM launch vehicle. It introduces the launch launcher, burn times and a summary of the vehicle and its major systems as well as manoeuvres of the upper stage, a detailed covering in particular detail the mission- launch event time line and trajectory de- specific equipment in the upper composite, scription, orbital ground-track as well mis- such as the payload adapter and separa- sion-specific analyses, e.g. upper stage tion system. The accommodation of the orientation angle to the sun, for the cus- payload is described accompanied by tomer. detailed clearance analyses, the fairing venting analysis, fairing jettison analysis, The resulting trajectory is then used as as well as measures for electrostatic dis- input data for various analyses such as charge control. orbit dispersion, loads, thermal, separation sequence and telemetry/ ground station 8.3 Trajectory and Mission coverage. Sequence (Book 4) The results of the Critical Design Review This book describes the mission timeline in provide the final flight data including: detail starting from the countdown se- • The flight event sequence for the on- quence to separation and upper stage orbit board computer removal/ de-orbiting. • The guidance parameters for the on- In order to perform these analyses, the board computer Customer is requested to submit the following detailed data in addition to the data 8.4 Dynamics of Spacecraft contained within the launch services con- Separation (Book 5) tract. This includes: This particular study provides a detailed • spacecraft orbital parameters including assessment of the spacecraft dynamics required injection accuracy, after separation from the upper stage. This • constraints on the upper composite / includes calculation of the separation payload orientation during the coast velocity of the spacecraft relative to the phase portion of the flight such as upper stage and the angular velocities of those required for thermal manoeuvres, the spacecraft after separation. A Monte- Carlo type analysis is used to provide a • orientation of the spacecraft at the statistical basis for the results taking into moment of separation, and account the uncertainties of the interface characteristics, e. g. variations in spring

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pusher force and connector disconnection • Heat dissipation for all applicable force characteristics. Thus, the analysis modes of operation during the mission provides a confirmation of the interface phases covered design to meet the Customer's velocity and • Interface descriptions (areas of contact, tip-off rate requirements. The results also conductive and/or radiative properties) provide a confirmation of the overall separation strategy including collision free • Thermal requirements for the environ- separation for both near and long term ment to be maintained during integra- scenarios. tion, launch and flight

The final mission analysis repeats and The detailed requirements of the space- confirms the studies performed during the craft thermal model to be provided by the preliminary analysis for the latest configu- Customer are summarised in the ration data taking into account the actual EUROCKOT specification ESPE-0009. Rockot/Breeze-KM and payload parame- The preliminary thermal analysis must ters. Thus it enables EUROCKOT to prove compatibility of thermal requirements and environmental conditions during the • define the data to be used by the on- following phases or identify areas of con- board computer for the orbital phase cern where modifications have to be (manoeuvres, sequence) agreed upon for those phases: • predict the clearance between the • Operations within integration facilities separated elements in orbital flight to verify collision avoidance including • Transportation to the launch pad Breeze-KM orbit removal. • Spacecraft integration on the Rockot 8.5 Thermal Environment (Book launch vehicle 6, parts 1 and 2) • Integrated phase until launch

The thermal environment study is imple- • Ascent mented to show thermal compatibility throughout the mission. Book 6 part 1 • Aerothermal heating after fairing sepa- covers the ground operations phase ration whereas book 6 part 2 covers the launch • Coast phase and ascent phase up to separation. The Customer provides a thermal model of the The final analysis will update the thermal spacecraft containing: compatibility study for all actual launch vehicle and spacecraft parameters. • Description of the thermal nodes (heat capacities, mass type, etc.)

• Internal thermal couplings of nodes (conductive, radiative and convective)·

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is particularly true for the launch vehicle to 8.6 Dynamic Coupled Loads spacecraft compatibility analysis. A new Analysis (Book 7) updated payload model will usually be The goal of the coupled loads analysis used in each successive CLA simulation (CLA) is to evaluate the limit levels and the run. frequency response of low-frequency This analysis is based on a spacecraft dynamic loading of the payload below 100 dynamic model submitted by the customer Hz as part of the Integrated Launch as specified in detail in EUROCKOT Vehicle (ILV) in the key ascent events. specification ESPE-0008.

The CLA includes as many simulation runs The payload experiences the worst-case as are required for each specific launch dynamic loads during stage 1 flight vehicle to spacecraft integration project including the worst-case dynamic pressure phase: environment that results from aerodynamic In the launch vehicle to spacecraft disturbances and lasts about 60 s. Loads compatibility study phase, it shall be experienced during powered flight of either established whether the ILV mechanical stage 2 or upper stage are much lower and environment is compatible with the payload are therefore of no interest in terms of requirements. payload load capability.

In the PDR phase, Thus, three major events are considered in the CLA: • preliminary levels of payload dynamic loads shall be evaluated, • Lift-off

• requirements for the adapter system • Worst-case dynamic pressure qmax shall be finalised and • Stage 1 shutdown

• acceptable notching levels shall be Axial dynamic loading of the structure determined, if required. results from transient processes in the In the CDR phase, propulsion unit and can be predicted with a fairly high accuracy. This is evidenced by • the system-level test conditions for the telemetry obtained in earlier missions. As adapter system and the separation for radial loads, these are caused by system shall be finalised, random phenomena such as thrust differences between cluster engines at lift- • payload dynamic loads during ascent off or shutdown, the ground wind direction shall be predicted and and speed at the time the ILV leaves the • dynamic displacements and sufficient transportation and launch container, and clearance to payload fairing shall be the wind direction and speed at the time verified. the dynamic pressure attains its peak level. Therefore, the CLA is carried out for two More then one CLA simulation run may be mutually orthogonal directions of radial required in any single project phase. This

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disturbance when studying the above key • time variations of Craig—Bampton events. These directions are usually accelerations and displacements for selected to run in payload orientation each payload spacecraft. planes. The appendices to the CLA Report may The CLA time interval for each simulated also include supplementary data such as event shall be long enough to ensure a • time histories of the quantities reliable prediction of load levels under specified by the Customer, study. • shock spectra of the pre-specified After a description of the calculation quantities, and method and the covered load cases for each event and each radial disturbance • any other data as agreed upon with direction, the CLA Report includes EUROCKOT. • a table of maximum CoG accelerations for each spacecraft, 8.7 Spacecraft Cleanliness Control (Book 8) • a table of maximum interface loads fo each spacecraft, This study provides an assessment of how the cleanliness requirements for the cus- • tables of maximum values of the tomer's spacecraft are implemented includ- quantities generated by recovery ing methods, standards and measurement matrices, and methods. The following items are covered:

• acceleration time histories and CoG • Particle cleanliness referring to the shock spectra for each spacecraft. concentration of the particles in the air The summary results in this report include within the clean rooms and the payload summary tables of maximum CoG fairing. Also the particle concentration accelerations and interface loads for the for surfaces located close to spacecraft events investigated. For a payload having such as the payload fairing. a clampband interface, the maximum • Organic contamination (optional) con- interface line loads and the clampband sidering the concentration of the or- tensioning specifications are provided. ganic compounds in the air within the Attached to the CLA Report are ASCII files clean rooms and the payload fairing. containing Also the organic compound concentration for surfaces located close to • tables included in the CLA Report and spacecraft such as the payload fairing. specifying the maximum CoG accelerations, the maximum interface • Consideration and mitigation of the loads and the maximum values of the outgassing and offgassing by space- quantities generated by recovery craft dispenser and payload fairing matrices, and

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• Consideration and mitigation of poten- and an on-board transponder is described tial pyrotechnic contamination from the in some detail. fairing and separation system The measurement system description is • Consideration and mitigation of plume mainly concerned with the capabilities of contamination by retro-rockets during this system and measurements undertaken second stage separation as well as on ground and in flight. Among other Breeze-KM thrusters during orbit or atti- things, a list of the parameters measured tude manoeuvres especially during col- by the ground measurement facilities is lision avoidance manoeuvres after provided. This list includes temperature, separation. humidity and loads during the ground operations. For the flight phase, the pa- The standard cleanliness analysis is per- rameters monitored include pressure, formed in two phases. The preliminary temperatures, loads as well as separation contamination analysis must prove that confirmation signals. This thorough charac- accumulated contamination can be kept terisation of parameters during ground within the specified limits or identify areas operations and launch allows EUROCKOT of concern where improvements have to be to provide an extensive and thorough post agreed. The final analysis will confirm launch evaluation giving the Customer full contamination compatibility for all actual visibility as to whether the ICD require- launch vehicle and spacecraft parameters. ments have been met and to provide les- sons learned for future missions. 8.8 Measurement System (Book 9) 8.9 Electromagnetic

This book provides a detailed overview of Compatibility Study (Book 10) the measurement system which covers The preliminary electromagnetic compati- both the ground and the flight operations of bility (EMC) study allows EUROCKOT to the Rockot launch system. check the compatibility between frequen- The ground measurement facilities, which cies and their harmonics used by the make up part of the overall measurement launch vehicle, the ground stations and the system, support the acquisition of data spacecraft during launch operations and required during ground operations. The flight. This study is based upon the space- flight measurement system has two main craft frequency plan including intermediate functions, to provide tracking of the launch frequencies from 14 kHz to at least 20 GHz vehicle during ascent within visibility of the which has to be provided by the Customer. ground stations and to downlink important It also considers the impact of radiated telemetry information from the vehicle emission caused by spacecraft or launch during the whole flight. vehicle on RF communication capabilities.

The tracking system of the Rockot The Customer is also requested to submit launcher which uses ground radar stations parameters of radio-telemetric equipment operating simultaneously with the Rockot

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transmission and reception systems during 8.11 Ground Electrical Cabling, ground preparation, in flight and immedi- Power Supply and Interfaces ately after spacecraft deployment before with Spacecraft GSE the Rockot/Breeze-KM transmission and (Book 12) reception systems are switched off. The This book covers in detail the design solu- Customer also shall provide limits for emis- tions for the ground electrical cabling and sions and susceptibility regarding radiated interfaces of the Customer's ground sup- disturbances. In case of conflict, the study port equipment (GSE) as well as a sum- will include an analysis of possible solu- mary of the available power supplies for tions related either to the launch vehicle or customer equipment. Specifically, this to the spacecraft. describes cables and harnessing in the The final EMC study considers the actual Undertable Room 7, where the Customer's configuration of the launch vehicle and GSE is located, as well as the test steps spacecraft. The study involves the exami- and check-out procedures used to verify nation of possible spurious emission fre- the correct installation and functioning of quencies and the susceptible frequencies these circuits. Furthermore, a detailed of the receivers. description of the available power supplies including uninterruptible power supplies is 8.10 Onboard Electrical Interface given. of Spacecraft and Launch Vehicle (Book 11) 8.12 Launch Base Operations Support (Book 14 part 1, 2, 3) This book covers in detail the configuration of the ground electrical cabling designed to Book 14 provides a summary of the agreed provide interfaces between the Customer's services necessary to support the Cus- electrical ground support equipment tomer launch site operations as provided (EGSE) on the one hand, and the space- from the EUROCKOT / KSRC side. This craft at the integration facility and the covers the responsibilities and support as launch site on the other. The extent of and well as the agreed processing schedule in procedures for electrical check-outs of the such areas as spacecraft transportation ground cabling are specified. The available within the launch site, off-loading opera- power supply systems are described to- tions of the Customer spacecraft container gether with the types, quantities and the and equipment at the facility, support for locations of outlets for connecting the standalone spacecraft processing, ground Customer’s EGSE at the integration facility operations support including support or the launch site. equipment such as boom lift, access platforms, fork lift etc. as well as safety aspects including definition of hazardous operations, crane operations, spacecraft fuelling support services and equipment. A successful conclusion of these aspects in the critical design review leads to the es-

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tablishment and release of the Joint Opera- 8.15 Social services (Book 16) tions Plan which then becomes the working document at the launch site. This book provides a summary of an important but often overlooked part of the launch campaign, i. e. the comfort and 8.13 Reliability (Book 15, part 1) welfare of customer and contractor per- This book provides an overview of the sonnel during their stay at the launch site. measures to ensure quality and reliability This book describes the city infrastructure, of the launch process according to Russian personnel transportation services available Federation standards. Specifically this within the launch site, the hotel and ameni- provides a description of the qualify assur- ties such as laundry services, satellite ance process used at KSRC including the television, medical services, dining facilities procedures, practices and testing methods etc. Therefore, the customer has the pos- used to verify this. Furthermore theoretical sibility to express their agreement with reliability figures for various Rockot/ these arrangements or to discuss other Breeze-KM subsystems are established to arrangements and services, e. g. special provide an overall reliability figure for the dietary requests like Asian food, at the complete Rockot/Breeze-KM launch sys- preliminary and critical design reviews. tem. Finally, a summary of the complete licensing process including the Customer 8.16 Communications support inputs and responsibilities is given. (Book 17)

This book reflects the services and infra- 8.14 ILV Components Quality structure available to support the Cus- Assurance (Book 15, part 2) tomer's communications requirements. This book describes the quality manage- This includes a description of the available ment system, namely, responsibilities, the telephone and facsimile services, internet system’s documentation, personnel train- access via LAN or direct dial-up, mobile ing and continuous improvement. The plan communications through walkie-talkies, of measures to implement the ILV quality intra-launch site high data rate communica- requirements is presented. tions such as fibre optic and microwave transmission as well as the provision of a The existing KSRC quality assurance multitude of satellite television channels in procedures are described in detail includ- the hotel. Furthermore, a thorough descrip- ing the respective manuals, operation tion of the communications channels and guidelines and quality inspection proce- back-up services available at the Rockot dures. Mission Control Centre is given. It is emphasised that the Customer will be enabled to track all life cycle phases of 8.17 Security (Book 18) each procedure envisaged by the Quality This book provides a description of the Management System. jointly agreed security plan between the customer and EUROCKOT/KSRC and the

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launch site authorities. This aspect can be Archangel Talagi airport to EUROCKOT's particularly important when dealing with facilities in Plesetsk Cosmodrome. This customers with technology sensitive pay- includes definition of the transportation loads wherein their national governments timeline, definition of cargo and containers impose high security requirements on to be transported, customs clearance, spacecraft and equipment. EUROCKOT paperwork requirements and specific re- and KSRC have gained extensive experi- sponsibilities, the transportation route and ence in this area especially in meeting the timing, interfaces to transport equipment, strict standards imposed for the launching such as lorries, rail wagons and lifting of payloads from the USA, wherein round devices, contingency planning for off nomi- the clock surveillance and restricted ac- nal situations, e.g. delays as well as ship- cess must be ensured around the satellite ping returnable equipment after conclusion and equipment. of the launch etc. It also provides an overview of the available transportation meth- 8.18 Transportation from Port-of- ods for personnel intending to travel to the Entry to the Launch Site launch site. The review of this book during Facilities (Book 19) the preliminary and critical design review allows the customer sufficient time to fine This book provides a thorough description tune this important aspect of the launch of the agreed transportation plan for the campaign such that a smooth and suc- customer spacecraft and equipment from cessful transportation of the spacecraft and its arrival at the Russian port-of-entry its equipment is assured.

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Chapter 9 Safety

Table of Contents

9 Safety ...... 9-1 9.1 Introduction...... 9-1 9.2 Submission Procedure ...... 9-1 9.2.1 Phase I Safety Submission...... 9-1 9.2.2 Phase II Safety Submission...... 9-3 9.2.3 Phase III Safety Submission...... 9-3 9.3 Safety Submission Contents and Requirements...... 9-3 9.3.1 Release of Safety Statements ...... 9-3 9.3.2 Final Date for Submission...... 9-3 9.3.3 Applicability...... 9-3 9.3.4 Identification of Statements ...... 9-4 9.3.5 Spacecraft Safety Data Package Contents ...... 9-4 9.3.6 Hazardous Systems...... 9-4 9.3.7 Guidelines for Safety Analyses...... 9-4 9.3.7.1 Overall Assessment of Risk and Severity...... 9-4 9.3.7.2 Probability of Hazards...... 9-5 9.3.7.3 Prevention of Hazards ...... 9-5 9.3.7.4 Reference Documents ...... 9-5 9.4 Non-compliance with Safety Requirements / Waivers ...... 9-5 9.5 Summary ...... 9-6

List of Figures

Figure 9-1 Safety submissions and assessment...... 9-2

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9 Safety time to take into account design constraints and measures necessary to meet the regulations and reduce the impact of costly 9.1 Introduction design changes late in the project. Figure 9-1 provides a schematic view of the sub- EUROCKOT is responsible for ensuring missions process. the compliance of the spacecraft, ground support equipment and launch site opera- It should be noted that the phased safety tions with the requirements of the stan- submission procedure described in the dards of the Russian Federation and of following sections is a generic description countries where the spacecraft and support for a spacecraft under development. For equipment are developed. For more de- existing spacecraft designs, the safety tailed information refer to the EUROCKOT submission process can be streamlined. Safety Handbook EHB-0004. The purpose of these regulations is to ensure the safety 9.2.1 Phase I Safety Submission of the environment, population, service personnel, ground equipment and facilities The spacecraft designer or Customer and the Rockot/Breeze-KM launch vehicle. prepares a file containing all planning documents for hazardous systems. The file It is the responsibility of the Customer and shall contain a description of the hazard- the spacecraft designer to ensure compli- ous systems and a reply to a hazardous ance with the safety requirements in the items check list supplied by EUROCKOT. design of the spacecraft, ground support A detailed check list of potential hazardous equipment and applicable processes. items can be found in the EUROCKOT KSRC will review each submission and Safety Handbook EHB-0004. issue an assessment report that will be subject to approval by EUROCKOT. A The document shall cover all safety-related spacecraft safety certificate and a safety activities such as component choice, safety data package approved by EUROCKOT and warning devices, risk analysis for are pre-requisites for obtaining a launch catastrophic events and in general all data licence from KSRC. enabling the risk level to be evaluated.

EUROCKOT will study this submission, 9.2 Submission Procedure classify the hazardous systems described To ensure early identification of the con- and declare any special requirements straints of the safety requirements upon imposed by the Flight/Ground Safety de- the spacecraft, support equipment design partments. EUROCKOT will compile a list and operations, the safety submissions are of potential sources of hazards in accor- split into three phases with the initial phase dance with the list provided in EHB-0004. I undertaken as soon as possible after Based on this list, EUROCKOT will compile contract signature. This allows the space- a list of reports on SC/GSE hazard analy- craft contractor and Customer sufficient ses.

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Customer Submissions EUROCKOT Reply Передача документов Заключение Заказчиком «Еврокот»

Phase I: Description of SC Systems (incl. Hazardous Systems); Preliminary Risk Analysis EUROCKOT (KSRC) Classification of Hazardous Systems; Compiling List of Reports on Risk Analyses Phase II: Description of Verification Methods (with References to Design and Operational Documents, Qual/Acceptance Docs) for Hazardous Systems; EUROCKOT Development of (KSRC) SC Processing Site Assessment Report for Phase Requirements II Submissions

Phase III: EUROCKOT Final Verification incl. (KSRC) Verification/Operation Procedures Approval of SC Safety for Hazardous Systems; Certificate Development of SC Safety Certificate KSRC Procurement of Launch License

Figure 9-1 Safety submissions and assessment.

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EUROCKOT’s replies to the submissions. 9.2.2 Phase II Safety Submission In addition, the spacecraft designer or The spacecraft designer or Customer shall Customer shall submit the final verification submit the hazardous systems manufactur- plan and operations procedures for sys- ing, qualification and acceptance docu- tems described as hazardous. mentation as soon as it becomes available. It must satisfy the requirements laid down 9.3 Safety Submission Contents by EUROCKOT at the end of phase I. This and Requirements documentation states the requirements for spacecraft integration facility equipment A detailed description of the format, con- and operations to be used during the tents and requirements for the safety sub- launch campaign and all other documents missions is given in the EUROCKOT required by EUROCKOT during phase I Safety Handbook EHB-0004. As a mini- and phase II submissions. It also defines mum, the format and data described below the policy for checking and operating all must be presented at the safety review of systems classified as hazardous. phase III, but no later than six months prior to launch. The final statement shall take EUROCKOT checks that the documenta- into account the responses from tion supplied in phase II complies with the EUROCKOT to the safety submissions requirements specified in phase I, states its made in phases I and II. intentions concerning verification of systems classified as hazardous and defines 9.3.1 Release of Safety Statements the draft procedure to be applied during spacecraft activities. The safety statement (safety certificate) is the official document from the spacecraft Phase II hazard analysis reports estab- designer or Customer and is signed by lished by the Customer shall describe with responsible officers of the spacecraft sufficient degree of detail: designer or Customer, e. g. by the Project • Hazard prevention measures Manager/Chief Designer of Spacecraft, Department Manager. • Methods of verification of hazard prevention measures with references to 9.3.2 Final Date for Submission drawings, manuals, etc. The final safety statement shall be submit- • National or agency safety standards ted to EUROCKOT not later than five with which the spacecraft complies months before the spacecraft launch.

9.2.3 Phase III Safety Submission 9.3.3 Applicability

The final safety submission must result in a A safety statement shall be presented for statement accompanied by a data package the following phases of operation: encompassing the complete results arising from phases I and II including

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• Operations with spacecraft and ground 9.3.7 Guidelines for Safety Analyses support equipment at the technical Analyses must be undertaken concerning complex and launch complex the safety of the spacecraft and the sup- • Flight of the spacecraft as a part of the port equipment during the following phases launch vehicle from moment of launch of operation: up to the spacecraft separation from • Operations with the spacecraft and the third stage. support equipment at the technical complex and launch complex 9.3.4 Identification of Statements • During the flight of the spacecraft as a Each safety statement is prepared in sepa- part of the launch vehicle, beginning rate lists and must include the following with the moment of launch up to information: spacecraft separation from the upper • Safety statement name and its desig- stage. nation number 9.3.7.1 Overall Assessment of Risk and • Name of the company which presented Severity this statement The results of the safety analyses of the • Name and post of the person who is spacecraft and the support equipment and responsible for it operations at the technical and launch complex are used to assess the overall • Date of submission safety of pre-launch operations. The format of this safety statement is The results of the safety analyses on the contained within annex 2 of the spacecraft during flight up to separation EUROCKOT Safety Handbook EHB0004. from the launch vehicle third stage are used to assess the overall safety of the

9.3.5 Spacecraft Safety Data Package flight phase of Rockot/Breeze-KM. Contents The safety of such phases is determined The data package which confirms the by the severity of the hazard impact classi- operational safety of the spacecraft and its fied with a severity of either catastrophic or support equipment shall be attached to the critical. safety statement. The package of data on safety concerning the spacecraft operation • Catastrophic severity is defined as the shall include data as described in Sections total loss of the launch vehicle and/or 9.3.6 and 9.3.7. spacecraft, ground facilities and/or support equipment. and/or severe injury or 9.3.6 Hazardous Systems loss of life to service personnel and/or severe damage to the environment and Please refer to the EUROCKOT Safety population. Handbook EHB-0004 for a more detailed list and description of hazardous systems.

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• Critical severity is defined as a partial • Inhibits. Safety inhibits of spacecraft launch failure, aborted launch because and support equipment, concerning in- of the launch vehicle and/or, the space- advertent operation of systems: craft and/or support equipment failure − Inadvertent operation leading to a ca- and non-fatal injury to service person- tastrophic failure must be inhibited by at nel. least three independent mechanical or electrical inhibits. In flight: − Inadvertent operation leading to a critical failure must be inhibited by at least • Catastrophic severity is defined as the two independent mechanical or electri- loss of the launch vehicle and/or the cal inhibits. spacecraft. 9.3.7.4 Reference Documents • Critical severity is defined as an incom- plete achievement of the mission goal, In the results of the analysis of the space- i. e. reduced mission effectiveness. craft or support equipment, there should be references to design and operational docu- 9.3.7.2 Probability of Hazards mentation, to procedures on equipment and system level, to existing statistics for The potential threat of a hazard must be previous spacecraft/equipment, or to verifi- evaluated qualitatively by classification of cation by similarity. In analyses of emer- the probabilities of such events occurring gency cases in which the spacecraft or into categories ranging from “High” to equipment endangers the launch or techni- “Remote”. cal complex or the launch vehicle during flight, any documents used as references 9.3.7.3 Prevention of Hazards in the assessment must be mentioned. In the event of any of the above mentioned In the safety analyses of the spacecraft documents having classified status and and support equipment, measures con- being unable to be released, a non- cerning prevention of hazards via the classified version of the document should spacecraft design, technology and opera- be provided. tional barriers must be shown.

The requirements for spacecraft safety 9.4 Non-compliance with Safety must include the following: Requirements / Waivers

• Design characteristics. Safety of EUROCKOT and KSRC shall identify the spacecraft and support equipment via non-compliances, if any, of the spacecraft design measures and material charac- and its GSE with the safety provisions of teristics such as: EHB-0004 as early as possible. The Cus- − Strength coefficients tomer shall document any non-compliance − Sealings/couplings structure and quality in form of a report on spacecraft non- of umbilical connectors compliance with the safety provisions of − Design and layout of the cable network EHB-0004 and submit this report to insulation EUROCKOT and to KSRC safety experts

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for review and provision of their recom- 9.5 Summary mendations. These recommendations shall be taken into account and implemented It must be shown as a result of the safety prior to spacecraft safety certification. The submissions described above that the non-compliance reports shall demonstrate spacecraft and its support equipment have that a particular provision or provisions in been subjected to analyses and tests EHB-0004 cannot be complied with and yet which confirm their compatibility with the the respective operational and/or manage- Rockot/Breeze-KM launch system for all rial measures introduced in connection with phases from launch preparation, through the non-compliance or non-compliances the launch and ascent phase and up to under review will minimize the risk of the spacecraft separation. The final safety associated hazards. Each report shall be submission (phase III) shall be presented signed by the spacecraft designer’s execu- to EUROCKOT not later than five months tives whose positions are higher than those prior to launch for approval. of the signatories to the hazard reports. It should be noted that EUROCKOT will The spacecraft safety statement shall be work actively with and assist the spacecraft signed at a higher managerial level if the contractor or Customer to help them meet spacecraft fails to comply with one or more the safety regulations. For this purpose, provisions in EHB-0004. EUROCKOT will interact with the Cus- tomer very early on in the mission integration process (phase I submission) to ensure that no surprises occur at a late date.

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Chapter 10 Plesetsk Cosmodrome

Table of Contents

10 Plesetsk Cosmodrome ...... 10-1

10.1 General Description...... 10-1 10.1.1 Climatic Conditions...... 10-3 10.2 Logistics ...... 10-3 10.2.1 Spacecraft and Hardware Transport ...... 10-3 10.2.2 Transportation Requirements...... 10-4 10.2.3 Transport Environments ...... 10-5 10.2.3.1 Environmentally Controlled Transport of Spacecraft during Launch Campaign...... 10-5 10.2.4 Spacecraft Team Transport to Plesetsk Cosmodrome...... 10-5 10.2.4.1 Charter Aircraft...... 10-5 10.2.4.2 Scheduled Flights...... 10-6 10.2.4.3 Rail Transfer to Plesetsk Cosmodrome ...... 10-6 10.2.5 Customer Team Transport at the Launch Site ...... 10-6 10.3 Communications...... 10-6 10.3.1 Phone Lines ...... 10-7 10.3.2 Data Lines ...... 10-7 10.3.3 Mobile Radios...... 10-8 10.3.4 CCTV...... 10-8 10.3.5 Entertainment TV ...... 10-9 10.4 Ground Facilities ...... 10-9 10.4.1 The Integration Facility MIK...... 10-9 10.4.1.1 General Hall ...... 10-12 10.4.1.2 Clean Room Bay ...... 10-12 10.4.1.2.1 Airlock ...... 10-12 10.4.1.2.2 Upper Composite Integration Area...... 10-12 10.4.1.2.3 Spacecraft Processing Area...... 10-14 10.4.1.3 EGSE Rooms...... 10-15 10.4.1.4 Customer's Office Area ...... 10-16 10.4.1.5 Handling and Hoisting Equipment in MIK ...... 10-16 10.4.1.6 Power Supply of MIK...... 10-16 10.4.2 The Rockot Launch Complex ...... 10-19 10.4.2.1 Launch Pad ...... 10-20 10.4.2.2 Undertable Rooms ...... 10-21 10.4.2.3 Power Supply of Launch Complex...... 10-21 10.4.2.4 Air Conditioning of the Spacecraft at the Launch Pad...... 10-23 10.4.3 The Mission Control Centre...... 10-23 10.5 Launch Campaign ...... 10-24

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10.5.1 Responsibilities and Operational Organisation...... 10-24 10.5.2 Planning ...... 10-25 10.5.3 Procedures and Logbook of Works ...... 10-25 10.5.4 Training / Briefings ...... 10-26 10.5.5 Security and Access Control ...... 10-26 10.5.6 Safety ...... 10-26 10.5.7 Launch Campaign Operations...... 10-26 10.5.7.1 Launch Vehicle Operations in MIK and at Launch Complex ...... 10-28 10.5.7.2 Spacecraft Operations ...... 10-28 10.5.7.3 Combined Operations in MIK...... 10-28 10.5.7.4 Combined Operations at the Launch Pad...... 10-29 10.5.8 Launch Day ...... 10-30 10.5.9 Abort Re-Cycle/ Return-to-Base Operation...... 10-31 10.6 Accommodation and Leisure Activities...... 10-31 10.7 Medical Care ...... 10-32

List of Figures

Figure 10-1 Geographical location of the Plesetsk Cosmodrome...... 10-2 Figure 10-2 Plesetsk Cosmodrome layout...... 10-2 Figure 10-3 Communications links...... 10-7 Figure 10-4 Layout of the integration facility MIK...... 10-11 Figure 10-5 Assembly stand with platforms in upper composite integration room...... 10-13 Figure 10-6 Scheme of removable fuelling platform...... 10-15 Figure 10-7 Customer office area in MIK...... 10-17 Figure 10-8 Launch complex for Rockot...... 10-20 Figure 10-9 Rockot launch pad...... 10-22 Figure 10-10 Undertable Room 7 for Customer's use...... 10-23 Figure 10-11 Schedule of operations...... 10-27 Figure 10-12 Flow of combined operations in MIK...... 10-29 Figure 10-13 Launch pad operations...... 10-30

List of Tables

Table 10-1 Characteristics of airports for shipping of spacecraft containers, related GSE and personnel transfer...... 10-4 Table 10-2 Specification of 208/120 V 60 Hz AC processing facility power supply system...... 10-18 Table 10-3 Specification of 380/220 V 50 Hz AC processing facility uninterrupted power supply system...... 10-19

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10 Plesetsk • The launch complex comprising: − Ground support facilities including Cosmodrome service tower, stationary mast, air conditioning system, fuelling system for stage 1 and stage 2, and equipment for launch preparation 10.1 General Description − Payload EGSE rooms in the undertable area The Rockot Plesetsk launch site is located at 62.7° N latitude and 40.3° E longitude, • Mission control centre (MCC) in Mirny, about 800 km north-east of Moscow and which provides: 200 km south of Archangel (Figure 10-1). − Accommodation for the Customer during launch The Plesetsk Cosmodrome was founded in − Customer console seating 1963 as a test range for launchers. Since − Countdown 1967 several international programmes − Data display and transmission of with the participation of France, Germany, launch information Great Britain and the USA have been − Operational communications launched from Plesetsk. Plesetsk Cos- − Room for VIP/Catering modrome conducted over one third of all • Integration facility MIK with launches in the world with a total number − General hall for offloading/ loading, of over 1500, including launches for some container cleaning and storage 30 Western and Asian satellites. Plesetsk − Clean room bay with upper composite Cosmodrome covers an area of 1752 km² integration room and processing room and includes the non-civilian Pero Airport for spacecraft integration, testing and for spacecraft fuelling and a railway station, the town of Mirny, − Administrative area with offices and LOX and LN2 plant, ground tracking sta- monitoring room tions, integration and technical facilities − EGSE and fuelling control rooms and launch pads. The main layout of the − Capability for environmentally con- Cosmodrome is shown in Figure 10-2. trolled spacecraft battery storage

The facilities used for the Rockot launch • Helicopter landing pad are: • Fuelling facility for the Breeze-KM upper stage. • Airport • Hotel Rockot in Mirny All the facilities used for the Rockot launch are linked by rail and road. The Rockot dedicated launch pad is adjacent and the helicopter landing pad is in the vicinity of the MIK (Figure 10-2).

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Figure 10-1 Geographical location of the Plesetsk Cosmodrome.

2 9 1 N

1 10

11 6

5 4

1 Tracking Station 1 Ground tracking station 8 2 River Emtsa 2 River3 RailwayEmtsa Head 3 Railhead 4 Pero Airport 35 km Integration Facility MIK – Fuelling Facility 7 4 Pero5 Areaairport 32T Technical Complex MIK Distances:6 km Integraton Facility MIK – Launch Pad 5 Area6 Fuelling32T technical Facility complex MIK 70035 m km Integration MIK - Facilityfuelling MIK facility - Helipad 6 Fuelling7 Area facility 133 ROC KOT Launch Complex 1006 km m Launch MIK Pad - launch– Blockhouse complex 7 Area 133 Rockot launch complex 700 m MIK - helicopter landing pad 8 Helicopter landing pad 18 km Launch pad - ground tracking station 9 Town of Mirny 45 km Pero airport - MIK Road 10 Hotel Rockot 20 km Pero airport - Town of Mirny Rail 11 Plesetskaya railway station 42 km MIK - Mirny

Figure 10-2 Plesetsk Cosmodrome layout.

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10.1.1 Climatic Conditions The Customer will also be fully supported by EUROCKOT and KSRC during the The climatic conditions in Plesetsk are customs clearance process at the port-of- continental, with the following characteris- entry. tics: Table 10-1 gives an overview of the char- • Minimum winter air temperature -38°C acteristics of Archangel Talagi and other • Maximum summer air temperature airports potentially used. Archangel Talagi +33°C is able to handle the largest Russian cargo aircraft, the AN-124, which is generally the • Average annual precipitation amount workhorse of the space industry for large 398 mm transports (Figure 1-4). This port-of-entry is usually used by Customers for the import 10.2 Logistics of their spacecraft and equipment to the Russian Federation. The main goal of the logistics items described in the following is to provide an The spacecraft and equipment transport to overview of how the transportation, espe- the EUROCKOT facilities is described cially of the spacecraft and related equip- below. An approximate timeline for a Cus- ment as well as customer personnel to tomer from Asia is provided as an example Plesetsk Cosmodrome can be performed. for preliminary planning purposes. Aircraft The transport logistics described herein are arriving directly from Europe and the USA all qualified and operational and have been will have different arrival times due to their used by all EUROCKOT Customer's to different time zones. date. Although not explicitly described, an identical but reverse transportation route • Day one morning: arrival of the trans- will be used for return of spacecraft equip- port aircraft (e. g. AN-124) at the Rus- ment and personnel after conclusion of the sian port-of-entry Talagi airport in Arch- launch campaign. angel. • Day one morning: off-loading and customs clearance in Talagi. 10.2.1 Spacecraft and Hardware Transport • Day one noon: loading of equipment onto EUROCKOT / KSRC supplied As a baseline, EUROCKOT's transporta- lorries and transfer at low speed to tion responsibilities begin upon off-loading railhead over an approximate distance of the Customer spacecraft and equipment of 10 km (Figure 1-5). at the Russian port-of-entry, Talagi airport in Archangel. EUROCKOT supported by • Day one afternoon to early evening: KSRC and its Russian subcontractors will transfer and securing of spacecraft and provide a team and equipment to conduct equipment from lorries to railway wag- off-loading and transfer activities from the ons (Figure 1-6, Figure 1-7). aircraft all the way to EUROCKOT's proc- • Day one evening: departure of railway essing facilities in Plesetsk Cosmodrome. wagons to Plesetsk train station with a

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travel distance of approximately 200 subsequent off-loading of equipment km. As the public railway system is (Figure 1-8). used on this route, the transportation is An optional transportation path for space- conducted during the night in order to craft fitting into a standard 20 foot con- allow low travel speeds. tainer or smaller is to fly to Pero airport at • Day two morning: arrival of train at Plesetsk Cosmodrome using an IL-76 after Plesetsk train station and transfer to in- an intermediate customs stop in Archangel ternal cosmodrome railway network. Talagi airport. • Day two morning: transfer on the cosmodrome internal rail network to the EUROCKOT processing facilities and

Parameters Sheremetyevo/ Pulkovo, Talagi, Archangel Pero, Plesetsk Domodedovo St. Petersburg Cosmodrome* Moscow

Status of airport International International International Non-civilian 3699 m Runway length 3780 m 2500 m 2000 m 3794 m Surface solidity, No constraints No constraints 44 11 PCN Types of AN-124, restrictions may IL-76, AN-72, AN- aircraft to be All types All types apply to some other wide 12, AN-24, YAK-42, accommodated body aircraft types Yak-40 Landing III A II II I category Role of airport for Port of entry back- Port of entry For small spacecraft Nominal Port of entry spacecraft shipping up back-up only * Pero Airport of Plesetsk Cosmodrome is operated and controlled by the Russian Space Forces (RSF) and can process civilian aircraft only if cleared by the Russian Ministry of Defence (MOD) general staff and possi- bly subject to special navigator availability on board. Table 10-1 Characteristics of airports for shipping of spacecraft containers, related GSE and personnel transfer.

10.2.2 Transportation Requirements reviews. These requirements are transferred to the Joint Operations Plan. The The basic transportation requirements for requirements shall cover the following at spacecraft and related equipment are least: described in detail in the Customer's response to the spacecraft questionnaire • Container handling and storage re- data sheet (chapter 12), sent at the begin- quirements ning of the project to customers and re- • Number, dimensions, weight, centre of viewed during the spacecraft to launch mass, and material of all containers vehicle preliminary and critical design

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• Container grounding requirements assembly and reloading on the transportation unit. The characteristics of conditioned • Necessity of immediate container air can be taken from chapter 5.2.3. transfer or possibility of intermediate storage upon arrival 10.2.4 Spacecraft Team Transport to Plesetsk Cosmodrome • List of hazardous materials and their international codes Customer personnel transfer from the • Maximum allowable duration for inter- Russian port-of-entry to Plesetsk Cos- ruption of container power supply and modrome, as well as the transfer from maximum/minimum allowable ambient hotel to launch site facilities, will be ar- temperatures during periods of cargo ranged and supported by EUROCKOT/ transhipment KSRC. Upon request, quotes for air / rail transportation fees will be provided. 10.2.3 Transport Environments For entry to Russia, EUROCKOT will sup- The spacecraft and its equipment will be port the Customer’s team in all necessary subjected to mechanical and thermal envi- formalities for customs clearance and in ronments during their transportation by air the obtaining of visas for the spacecraft and on the ground as well as during team. The entry to Russia and departure ground handling. In section 5, the worst from Russia will also be assisted by case transportation and handling loads are EUROCKOT. described. Electrical power can be sup- Normally, the spacecraft team will arrive at plied to the Customer for environmental Sheremetyevo or Domodedovo Interna- control if the spacecraft container does not tional Airport, Moscow. The ongoing possi- have its own power. Additionally, the re- bilities of transfer to Plesetsk Cosmodrome sults of transportation load measurements are described in the following chapters. for rail and road transport are included.

10.2.4.1 Charter Aircraft 10.2.3.1 Environmentally Controlled Transport of Spacecraft during Launch Pero Airport at Plesetsk Cosmodrome has Campaign limitations on the aircraft size to be ac- For the transport of spacecraft as part of commodated (Table 10-2). For the trans- the launch vehicle upper composite from port of the Customer’s personnel, a YAK- the payload processing facility (MIK) to the 40 aircraft will normally be used. The YAK- launch pad, a mobile thermal conditioning 40 can be provided in an economy class system is available (Figure 1-26). configuration with 22 seats or in a configuration with 5 business class seats and 10 In order to maintain the required tempera- seats of economy class. The flight duration ture, moisture and cleanliness conditions from Moscow to Plesetsk is approximately under the fairing in the vicinity of the 2 hours. spacecraft, the thermal conditioning unit is connected to the upper composite after its

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10.2.4.2 Scheduled Flights the military restricted area of the Cos- modrome will be organised by The nearest airport to Plesetsk Cos- EUROCKOT. modrome for scheduled flights is to Talagi airport in Archangel. The flight route Mos- 10.2.5 Customer Team Transport at the cow - Archangel with a flight duration of Launch Site approximately 1 hour 50 minutes is a regular scheduled route. EUROCKOT can At the launch site, buses, minivans and, if arrange a special transfer from the airport necessary, lorries will be made available to to Plesetsk Cosmodrome by car taking transport the management team, technical approximately 5 hours. support and security personnel to the work area during spacecraft processing and launch operations according to the daily 10.2.4.3 Rail Transfer to Plesetsk Cosmodrome working schedule. The distance between Hotel Rockot, where the Customer is ac- A convenient alternative and non-weather commodated, and the technical complex is constrained option for travelling to Plesetsk 42 km. The transfer time by bus is ap- is by rail. The rail transfer from Moscow to proximately 45 minutes. Plesetskaya, which is 3 km south of Mirny, is via overnight train and takes 18 hours. This transportation route is especially 10.3 Communications suitable and cost-effective for transportation of small groups of personnel to and Reliable and independent national and from the launch site. Please note that the international communication services are overall travel time for the customer to the provided by the telecommunication system launch site from their foreign facilities is not installed in Plesetsk Cosmodrome. All much longer than via other routes. The positions in the processing facility MIK, travellers sleep in the train rather than launch pad, mission control centre (MCC) staying overnight in Moscow and arrive the and Hotel Rockot are interconnected by next day, as they would if they flew by microwave or fiber-optic lines that are charter aircraft. The sleeping compart- connected to international lines through ments are of good standard and are com- Moscow via satellite (Figure 10-3). The fortable and clean. Booking of sleeping following types of optional telecommunica- compartments in the train can be arranged tion services are available upon request: through EUROCKOT as well as a transfer • Local and international direct dial (IDD) to a domestic airport or railway station in phone lines Moscow. • Data lines If customers arrive by plane in Archangel, • LAN a transfer to Plesetsk can also be arranged by train taking approximately 5 hours. • Mobile radios (IDD) • CCTV The transfer from Plesetskaya station to Hotel Rockot in Mirny, which is located in • Inmarsat

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• Internet access and mail server • Entertainment TV in the hotel • Various types of telecommunications support at the MCC (Section 10.4.4)

Figure 10-3 Communications links.

10.3.1 Phone Lines phones. Access to mobile radios is possible within the phone network. The national and international phone lines support links between launch site facilities, the Mission Control Centre and the hotel. 10.3.2 Data Lines International calls can be placed from any phone line. Multi-purpose RJ 45 jacks are In order to support data transmission at the used to connect either digital or analogue launch site, inter-area data lines connecting the integration facility MIK, the launch

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complex and the hotel as well as interna- fibre optic lines with a data rate of up to tional data lines are available. The launch 10 MBits/s. site inter-area data lines comprise: • Analogue interface for modem-based 10.3.3 Mobile Radios transmissions at up to 19.2 kbps Mobile radio links will be available at the • ISDN interfaces for data rates of up to MIK, at the launch site, in the Customer 128 kbps accommodation area and also along the • V.35 interfaces for data rates of up to routes interconnecting any two of these 64 kbps locations. The mobile radio system operates in half duplex mode and supports • Data transmission over a multi-mode either conference sessions or point-to-point fibre optic cable between the Undert- links, with mobile radio access to or from able Room 7 and the integration facility the launch-base phone network being • LAN cable network in the MIK and possible. The required quantity of calling Hotel Rockot with RJ45 interface. To groups will be pre-programmed in each interconnect the MIK LAN segment and portable radio in use by the Customer. hotel LAN segment into a network, a 10.3.4 CCTV 2048 kbps channel is provided that runs through the MCC. Security video monitoring services will be • Dial-up access to Internet and a mail provided at the integration facility and the server of up to 30 users at a time with launch complex. The system enables V.90 protocol (56 kbps) from any ana- monitoring of the spacecraft and the Cus- logue phone interface tomer’s GSE with the image sent to the Customer’s rooms, taped and/or played • For access to Internet from the cus- back. The following video support is avail- tomer’s LAN, a network infrastructure able at the launch site: using Ethernet data lines is provided. Internet access from MIK and/or • Explosion-proof video cameras avail- Rockot Hotel via dedicated lines to able in all clean rooms with the camera Moscow ISP makes use of a satellite outputs delivered to the Customer’s of- channel with 64 Kbps to 2048 Kbps fice data rate. • Video monitors available in the Cus- • For spacecraft check-out while on the tomer’s office. Each video camera can launch pad an Ethernet data exchange be remotely controlled, i. e. panned, channel between the MIK and the tilted, focused and/or zoomed from this launch pad can be provided. The Cus- office tomer's EGSE LAN segment in the MIK • Video taping capabilities and on the launch pad can communicate via media converters with Ethernet • Video cameras available in the Undert- protocol operating through multimode able Room 7 to monitor the Customer’s GSE. The outputs from these cameras

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will be sent to the Customer offices in • Emergency exits with emergency the MIK. showers and eye wash facility • Office area 10.3.5 Entertainment TV

Four different TV channels in English, Access to the Customer area on the first German and/or other languages upon and second floor of MIK can be gained via request are available in each room of the a separate entrance with staircase or Rockot Hotel used by the Customer’s staff. escorted via the general MIK entrance. An additional staircase in the Customer's office area allows direct access to the foyer 10.4 Ground Facilities of the Customer's changing rooms. En- The facilities available to the customer for trance to the spacecraft processing area their spacecraft processing and launch and EGSE rooms, as well as to the upper campaign activities are described below. composite integration area in the case of joint operations, is possible without contact EUROCKOT's facilities at the launch site with military controlled areas. have been specially designed and con- structed to enable convenient implementa- At the launch complex, the Undertable tion of customer security measures. For Room 7 is dedicated to and under the instance many of the facilities can be security control of the Customer. placed under the sole control of the Cus- Video observation of the clean room area tomer and, therefore, under his security as well as the Undertable Room 7 can be control. For these areas, the Customer can performed from the Customer’s security implement access control procedures in office. During launch, the separate mission accordance with Customer state govern- control centre is used. mental regulations. Within the MIK, the Customer has its closed dedicated area under its sole security control. Within this 10.4.1 The Integration Facility MIK dedicated area the Customer can move without escort. This closed dedicated area The integration facility MIK (Figure 1-8, contains: Figure 10-4), internally designated building 130 in area 32T, is located in the south- • A spacecraft processing area for con- eastern part of the Cosmodrome at a dis- ducting autonomous spacecraft opera- tance of 6 km from the EUROCKOT launch tions and spacecraft fuelling complex and 42 km away from the town of Mirny. The spacecraft container and re- • EGSE rooms for support equipment lated equipment can be directly delivered installation to the MIK by helicopter. A helicopter land- • Change rooms, shower and rest rooms ing pad is located in the vicinity of the MIK. as well as an air shower The integration facility is intended for acceptance, storage, assembly and checking

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of boosters, Breeze-KM and fairing, accep- of 9150 × 7620 mm contains a system for tance of spacecraft, spacecraft operations containment and drainage of minor spills and assembly of the upper composite. It smaller than 1 litre of hydrazine propel- comprises: lants. The spacecraft will be placed on a special-purpose 3000 × 3000 mm support • General Hall with common work area adapter. Room 111A can be used as a • Clean room bay, certified cleanliness fuelling control room. A viewing window ISO Class 8 with between these rooms enables visual con- − hardware air-lock (54 m²) and person- tact. nel air-lock (101V) EUROCKOT can optionally provide a − clean room for the upper composite encapsulation (101B, 146 m²) fuelling service performed by an experi- − clean room for the spacecraft process- enced and qualified team. ing and fuelling (101A, 180 m²) The clean room bay is equipped with an • Two EGSE rooms (111 and 111A) oxygen content control system and a hy- • Administrative area with Customer drazine monitoring system. Critical values offices and rooms for remote control will automatically initiate alarm by acoustic and visual means. Fire protection in the Customers requiring clean room cleanli- integration facility is provided not only in ness conditions better than the given stan- the usual way but also by a water deluge dard should contact EUROCKOT for fur- system on the walls and by water curtains ther information. Fuelling of spacecraft can on the doors of the clean room bay. take place on a fuelling platform in Room 101A within the closed clean area de- The floor in the MIK is made of anti-static scribed above. This platform with an area and electrically conductive linoleum.

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Figure 10-4 Layout of the integration facility MIK.

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• Wall mounts for 120 V 60 Hz and 380 / 10.4.1.1 General Hall 220 V 50 Hz power supply

The common work area (Figure 10-4) with • 2 t overhead crane 500 m² in the general hall of the MIK is • Sensors, acoustic and visual warning dedicated for ground support equipment devices of oxygen control system unloading, unpacking/ packing and short- • Sensors, acoustic and visual warning time storage, booster preparation, autono- devices of hydrazine/ oxidiser control mous checks of Breeze-KM and fairing, system and electrical checks of the upper composite. For lifting operations, an overhead • Particle counter travelling crane with 30 tonnes lifting ca- • Grounding terminals pacity is installed. The general hall is not environmentally controlled but can be kept • Emergency lighting at reasonable temperatures. • Fire protection system • The size of the door leading from the 10.4.1.2 Clean Room Bay general hall to the airlock is 4.25 m wide x 11 m high The spacecraft processing as well as upper composite integration and encapsulation will take place in the closed clean 10.4.1.2.2 Upper Composite Integration Area area of the MIK (Figure 1-14) which is The upper composite integration area has certified to cleanliness ISO Class 8. The an area of 146 m² and is located between particular parameters of environmental the airlock and the spacecraft processing control can be taken from Section 5.2.2. area. The upper composite integration area Higher cleanliness levels can be provided is intended for the mating of the spacecraft as an option. with Breeze-KM and assembling of the upper composite (Figure 10-5). The 10.4.1.2.1 Airlock equipment of the upper composite integration area is stated below. The upper com- The airlock (room 101V) located at the posite integration area is accessible for the beginning of the clean room bay has a floor Customer’s personnel from the clean room area of 54 m² (9.4 m wide and 6 m long). changing area under the Customer's secu- An overview of the equipment of the airlock rity control after they have passed through is given below. The final cleaning of the the air shower. spacecraft container and associated equipment will be done here. The facility has two main doors for movement of spacecraft and equipment. The Equipment of the airlock: door to the airlock is 5 m wide x 11 m high, while the door to the spacecraft processing • Explosion-proof cameras for remote area is 4 m wide x 6.5 m high. control • Explosion-proof phones

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Figure 10-5 Assembly stand with platforms in upper composite integration room.

The upper composite integration room is • Emergency exit with emergency equipped with: shower and eye wash facility

• Explosion-proof cameras for remote • Sensors, acoustic and visual warning control devices of oxygen control system • Explosion-proof phones • Sensors, acoustic and visual warning devices of hydrazine/ oxidizer control • LAN-drops system • Wall mounts for single and three phase • Particle counter 208/120 V 60 Hz and 380/220 V 50 Hz power supply • Grounding terminals • 10 t overhead crane • Emergency lighting • Assembly stand shown in Figure 10-5 • Fire protection system

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10.4.1.2.3 Spacecraft Processing Area support operations involving high-pressure gases are standard services. The spacecraft processing area (room 101A), placed under Customer security A rescue team will be on duty throughout control, covers a total area of 180 m². This the fuelling operations. This team will area of the clean room bay is intended for include a fire engine, an ambulance and a spacecraft processing and comprises a 90 medevac car each fully manned. m² work area and a spacecraft fuelling Equipment of the spacecraft processing area with 90 m². The spacecraft processing room: area is accessible for the Customer’s personnel through a separate entrance • Explosion-proof cameras for remote from the clean room changing area under control the Customer's security control after they • Explosion-proof phones have passed through the air shower. Two emergency exits are installed in the north • LAN-drops and the east of the processing area, each • Wall mounts for 208/120 V 60 Hz and comprising two showers and one eyewash 380/220 V 50 Hz power supply facility. All operations in this spacecraft processing area can be monitored and • 10 t overhead crane recorded. Additionally, an explosion-proof • Removable fuelling platform with a window provides a view from EGSE room spillage containment and drainage col- 111A which can be used as a fuelling lection system control room, (section 10.4.1.3). • Two emergency exits with emergency The facility's main door for movement of showers and eye wash facilities spacecraft and equipment from the upper • Sensors, acoustic and optic warning composite integration area is 4 m wide x devices of oxygen control system 6.5 m high. • Sensors, acoustic and visual warning For spacecraft fuelling operations, devices of hydrazine / oxidizer control EUROCKOT can provide a removable system fuelling platform (Figure 10-6) with a sys- • Particle counter tem for containment and drainage of minor spills of not more than 1 litre and a • Grounding terminals spacecrafft support adapter. This fuelling • Emergency lighting platform will be located in the spacecraft processing area. The fuelling platform is • Fire protection/extinguishing system further designed to accommodate the • Industrial-grade compressed air supply fuelling equipment and propellant contain- • Industrial-grade water supply to wash ers. down spillages The provision of industrial-grade com- • Blast shield to support operations pressed air and technical water to wash involving high-pressure gases out spillage as well as a blast shield to

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A-A (not to scale)

Figure 10-6 Scheme of removable fuelling platform.

10.4.1.3 EGSE Rooms Room 111A is equipped with a viewing window to observe operations in the Adjacent to the clean area, an EGSE room spacecraft processing area. If fuelling of with an overall floor area of 100 m2 is the spacecraft is performed in the space- located. A wire-mesh bulkhead with a 2.5 m craft processing area, room 111A can be wide × 2.5 m high sliding door subdivides this used as a fuelling control room. room into room 111 with 60 m2 and room 111A with 40 m2. Room 111 has a door 3 m wide x 3.5 m high to transfer hardware from outside the These rooms can be used for accommoda- building. tion of the electrical equipment and fuelling equipment and can accommodate stations There is a 2.1 m wide x 2.1 m high door tooled up to support spacecraft processing. between room 111 and the spacecraft processing area. This door can be opened The Customer’s personnel can access for the time the GSE is being set up. How- room 111A from the office area on the ever, it will be airtight sealed during clean second floor. operations.

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If the spacecraft battery stored in room 111 fax capability, LAN- drops and a 60 Hz and needs cooling, the MOTEK battery cool 50 Hz power supply. Room 210 of the cabin built around a 20 foot sea shipping Customer’s office area provides direct container can be provided upon request. access to the changing rooms of the clean area by stairs. Room 209 provides video Equipment of EGSE rooms: monitoring capabilities for the TV cameras which are installed in the processing area • Cameras for remote control for security and safety needs. • Monitors (optional) Office equipment such as fax devices, • Phones electronic data processing system and • LAN-drops computer monitors, copy machines, overhead projectors etc. are available upon • Wall mounts for 208/120 V 60 Hz and request. 380/220 V 50 Hz power supply • Ground terminals 10.4.1.5 Handling and Hoisting Equipment in • Emergency lighting MIK • Common fire protection system The main handling and hoisting equipment in MIK comprises: • Optional optical and/or RF link terminals to the SC and/or Customer's • Boom lift EGSE at the Launch Pad. • Fork lift

10.4.1.4 Customer's Office Area • Rail car

The Customer’s offices are located on the • Trolley for adapter with spacecraft 2nd floor of the integration facility exten- • Mobile integration table sion in the vicinity of EUROCKOT and • Upper composite assembly stand KSRC offices (Figure 10-7). The second floor is above the EGSE rooms and clean room changing area. Access to the Cus- 10.4.1.6 Power Supply of MIK tomer’s office area is provided via the Uninterrupted power is supplied in all common MIK entrance or by a separate Customer dedicated areas in the MIK, staircase. The entrances to the Customer’s airlock and Upper Composite integration office area corridor are secured by lockable hall with 208/120 V at 60 Hz and 380/220V doors. at 50 Hz. The administrative area for the Customer The 208/120 V, 60 Hz Power Supply Sys- comprises seven offices (4 x 41 m² and 3 x tem (PSS) is a self-contained power supply 20 m²) dedicated to sole Customer use system set incorporating two diesel- (Figure 10-5). All rooms are equipped with generators and two uninterruptible power heating, smoke detectors and fire extin- sources (UPS). The PSS specifications are guishing systems and have a telephone / listed in Table 10-2.

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Figure 10-7 Customer office area in MIK.

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The power supply system can operate non- ers providing 100 kW power for not less stop for 30 days before it is turned off for than 10 minutes. maintenance work. In the case of failure of 50 Hz UPS power is also provided in the either a diesel generator (DG) or an unin- EGSE rooms as single or three phase terruptible power source (UPS), the other power supply to a total maximum rate of 30 DG or UPS will run at rated 100% power kVA. The specification of the 50 Hz UPS output. The uninterruptible power source power supply system is given in Table includes built-in battery-supported convert- 10-3.

Item Description

Rated power output 100 kV⋅A ( cos φ = 0.8) Power change under load (power consumers) From 0 to 100% of the rated output Line/phase output voltage 208/120 V Voltage waveform Continuous sinewave Current frequency 60 Hz ±1%, continuous with the load increasing from 0 to 100% Stable output voltage variation and falling back to 0 110% ≤ 20 min 125% ≤ 10 min Overload capacity at power output of 150% ≤ 1 min 200% ≤ 1 s Stable output frequency deviation ±0.1 % < 3% in static conditions Total harmonic distortion factor < 5% in dynamic conditions Storage battery capacity At least 10 min Voltage regulator response time < 5 ms Radio interference level Below "N" as per VDE 0875 Efficiency at rated load > 90% Protection to DIN 40050 Standard 1P21 Duration of system operation Continuous and uninterrupted for up to 30 days Three-phase, five-wire (A, B, C – phases, Type of system of current-carrying conductors N – neutral wire, PE – protective earth wire)

Table 10-2 Specification of 208/120 V 60 Hz AC processing facility power supply system.

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Item Description

Rated power output 30 kV⋅A (cos ϕ = 0.7) Power change under load (power consumers) From 0 to 100 % of the rated output Line/phase output voltage 380/220V Voltage waveform Continuous sinewave Current frequency 50 Hz Output voltage variation with continuous load varia- ±1% within period from 0 to 100 % and falling back to 0 tion Output voltage variation with stepwise load variation ±5% within period from 0 to 100 % and falling back to 0 Stable output frequency deviation ±0.1%, continuous Storage battery capacity At least 10 min Efficiency at rated load > 90 % Duration of system operation Continuous and uninterrupted for up to 30 days Capacity of storage battery Not less than 10 minutes Three-phase, five-wire (A, B, C – phases, Type of system of current-carrying conductors N – neutral wire, PE – protective earth wire)

Table 10-3 Specification of 380/220 V 50 Hz AC processing facility uninterrupted power supply system.

10.4.2 The Rockot Launch Complex The Rockot Launch Complex is situated in the north-east area of Plesetsk Cos- The launch complex as shown in Figure modrome, at a distance of 6 km from the 10-8 is dedicated to Rockot launches MIK. The railroad ends directly in front of exclusively for use by EUROCKOT. In the the launch table for loading and unloading course of rebuilding and renovating work the booster stages and upper composite. A for all Rockot dedicated facilities in the site plan of the Rockot launch complex is years 1999 and 2000, the Rockot launch shown in Figure 10-8. complex underwent a complete renewal, so that today all equipment is state-of-the- art.

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N Neutralisation Railroad Station Transformer

Fuel Storage Waste Deposit Lightning Road Protection Blockhouse

Service Channels Supply Control Tower Station (Utility Room) Water Storage

Rail Oxidiser System Storage Not to scale

Figure 10-8 Launch complex for Rockot.

10.4.2.1 Launch Pad with the Rockot booster. In the closed gate position, the service tower encloses the The launch pad includes a launch table Rockot to allow all-weather operations. The with launch equipment and a stationary service tower is equipped with a lift and mast, surrounded by the mobile service working platforms at several levels to tower during launch vehicle processing provide access to the launch vehicle ser- (Figure 10-9). vice areas. A special adapter frame serves The launch table is equipped with retract- for the transportation and launch container able supports for the alignment of the (TLC) erection, whereas an overhead launch vehicle and with metal gas deflec- travelling crane ensures zero impact mat- tors. The surface around the launch table ing of the upper composite with stage 2 of is covered with reinforced concrete. the Rockot booster unit.

The mobile service tower is designed for Retraction of the mobile service tower vertical integration of the Upper Composite occurs approximately 10 minutes before

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lift-off. At lift-off, the distance between the situated in room 213/111A of the Customer rolled-back service tower and the station- office area in the MIK. The Undertable ary column with Rockot in the TLC is ap- Room 7 is linked to room 213/111A via a proximately 50 m (Figure 1-23). fibre optic cable. The Customer shall provide the equipment for connecting the The stationary column is designed to fas- fibres. ten and hold the booster unit in the TLC at the moment of launch. Electrical cables, air The harness length from the Undertable ducts and fluid lines to the launch vehicle Room 7 to the spacecraft umbilical connec- are maintained via the stationary column. tors is approximately 80 metres. The stationary column also accommodates The equipment of the Undertable Room 7 the ground control equipment devices, the includes: launch vehicle azimuth positioning system and the remote control system. • Camera for remote control monitoring due to security reasons 10.4.2.2 Undertable Rooms • Phones

Undertable rooms contain the launch vehi- • Wall mounts for 208/120 V 60 Hz and cle EGSE and pre-launch processing 380/220 V 50 Hz power supply equipment, the fuelling system for booster • Fibre optic cable access stage 1 and stage 2 and the payload air conditioning system. For the Customer's 10.4.2.3 Power Supply of Launch Complex use, one undertable room, designated as Undertable Room 7 with an area of 26 m² The launch complex has its own independ- is provided (Figure 10-10). This Undertable ent UPS system. Power of 208/120 V at 60 Room 7, containing the equipment stated Hz with nominal output power of 100kVA below, is dedicated for the accommodation and 380/220V at 50 Hz with 10 kVA is of the spacecraft on-board battery trickle supplied in the Undertable Room 7. charging equipment and other items at the launch pad. Access to the undertable The launch complex power supply system rooms is possible except during launch is similar to that of the MIK (section vehicle fuelling operations, during final 10.4.1.6). The only difference is the rated countdown and launch. Monitoring equip- output power of 380/220V at 50 Hz equal ment for the spacecraft parameters can be to 10 kV⋅A.

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Figure 10-9 Rockot launch pad.

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Figure 10-10 Undertable Room 7 for Customer's use.

The air conditioning lines will only be dis- 10.4.2.4 Air Conditioning of the Spacecraft at connected during lift-off. In the case of a the Launch Pad launch cancellation the air conditioning system will be switched on again within 1 After encapsulation, air conditioning of the minute. The air conditioning system for the payload is provided during transportation of spacecraft and spacecraft batteries is the upper composite to the pad and up to compatible with ISO Class 8 cleanliness or 30 seconds before lift-off. Depending on optionally ISO Class 7 (chapter 5). the fairing design, purging of the spacecraft EUROCKOT will provide periodic monitor- batteries can also be performed, if re- ing of air temperature, humidity and flow of quired. Interruption of the payload air con- fairing air with a minimum interval of 15 ditioning occurs only for periods of no minutes during the time the spacecraft is longer than one hour each. on the launch pad. At the launch pad, the stationary air conditioning equipment is located under the launch table and the air duct passes 10.4.3 The Mission Control Centre through the stationary mast. The air duct The Mission Control Centre (Figure 1-26), and the air inlet of the fairing are con- located in the town of Mirny at a distance nected via a flexible tube.

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of approximately 40 km from the launch the viewing angle and solar exposure. pad, provides: In this mode, staging events are displayed and several upper stage per- • Local, interurban and international formance quantities are presented in telephone communications with the ca- numerical form pability of telephone conferencing • 3D motion of the item's CoG against • Video and audio reporting from the the earth background and several orbit launch site in real time parameters in numerical form • Countdown information • Generalised state vector including • Transmission and receive of the for- geodetic predicted position as well as mats in electronic and fax form actual orbit parameters, and the predicted as well as telemetered Breeze- • Operational communications with KM engine performance launch support services including backup communications, Inmarsat and • Live launch coverage public GSM The MCC is equipped with hardware and • Display and transmission of launch software which allows integration of any information in real time other information, coming e. g. from other monitoring facilities during and after • Go/ no-go system for the spacecraft spacecraft injection, into the set of data operator as well as go/ no-go display displayed. In addition, processed flight data panel and launch vehicle status. can be compressed and sent to any re- The following data can be displayed on mote user to be decompressed and dis- large size wall panels: played.

• CoG motions of stages 1 and 2, the 10.5 Launch Campaign upper stage and the fairing during powered flight in the ascent phase, 10.5.1 Responsibilities and Operational ranging data, key flight events such as Organisation staging and jettisoning of the fairing, major telemetry data down link and tip- EUROCKOT / KSRC is responsible for the off angles. preparation of the launch vehicle and • Sub-orbital unit track on an earth map combined operations. The Russian Space accompanied by the display of ranging Forces execute Rockot/Breeze-KM opera- data, key flight events such as upper tions including launch as subcontractors to stage burns, major telemetry data down EUROCKOT / KSRC. link timelines for ground telemetry sta- EUROCKOT/KSRC will conduct the instal- tions, and predicted or actual orbit pa- lation of the Customer’s spacecraft on the rameters adapter. Additional support shall be mutu- • Computer-generated presentation of ally agreed between EUROCKOT / KSRC spacecraft motion about its CoG, with and the Customer. The spacecraft prepara-

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tion will be performed under the responsi- working conditions for the Customer. Due bility of the Customer and its launch site to the parallel processing of spacecraft and team. launcher up to the joint operations prior to launch, these activities have to be coordi- During the launch campaign, a core of the nated to ensure the availability of neces- EUROCKOT team responsible for the sary equipment and support personnel and specific Customer mission will be present the accessibility of facilities, taking into at the range as day-to-day intermediaries account the security and safety matters of between Customer, KSRC and the range all parties involved. authority to coordinate the spacecraft launch site support requirements as well as During the launch campaign, a daily to accompany the Customer’s launch site schedule meeting will be held with the team in all matters. The EUROCKOT team participation of all parties involved, Cus- is supported by a KSRC team at the range. tomer, EUROCKOT/KSRC and attendees Both teams ensure undisturbed execution from the Russian Space Forces . The goal of all necessary operations until launch and of this meeting is to the fulfilment of spacecraft support requirements in accordance with the launch • communicate the status of the work site requirements. • identify issues that require immediate 10.5.2 Planning attention • define the schedule and coordinate Spacecraft launch site operations and the operations for the next day with a view relevant requirements will be specified in to the support personnel needed, ac- the Spacecraft Operations Plan (SOP) as cess to facilities, transportation needs, well as the responses to the spacecraft lunch times, questionnaire from EUROCKOT. These Customer-generated documents should • coordinate future joint operations and address all operational and logistical sup- • adapt the launch campaign schedule, if port requirements. necessary. All spacecraft activities and technical facilities will be controlled at the launch site 10.5.3 Procedures and Logbook of according to Joint Operations Plan (JOP) Works jointly established with EUROCKOT. Every process will have an approved pro- The JOP gives an overview of the space- cedure. These procedures will identify the craft operations and joint operations to be necessary equipment, personnel, docu- conducted at the launch site, and defines mentation and facility requirements in ground rules for all involved parties at the detail to complete the process. The related range. The JOP is established to define the launch site procedures will be carried out equipment and support needed at the under consideration of safety and security launch site for both spacecraft and joint regulations of the Russian Government operations in order to ensure undisturbed and the Customer state government. All procedures for joint operations have to be

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signed off by the Customer and 10.5.7 Launch Campaign Operations EUROCKOT / KSRC. The launch campaign operations, espe- All joint operations will be documented in a cially the spacecraft operations described logbook. The joint working steps are in the following, serve the purposes of documented in Russian and English lan- orientation. For a programme, the duration guage and have to be signed off by all of a launch campaign will be tailored to the parties upon completion of the work. Customer’s requirements. A final and detailed Launch Operations Schedule which includes a statement of the precise 10.5.4 Training / Briefings duration of all operations, will be established after definition of the Joint Opera- Training and briefings for the spacecraft tions Plan (JOP) together with the Cus- operations team will be performed before tomer. the start of the spacecraft operations. Such training and briefings comprise: A typical Customer launch campaign from the arrival of the spacecraft and related • Familiarisation with emergency evacua- equipment at the Cosmodrome until launch tion procedures and all alarms will last approximately 28 days (Figure • Communications equipment operations 10-11), up to three days needed for post- launch activities have to be added. A com- • Security requirements and briefings plete launch campaign, which also takes • Training to operate launch site specific the launch vehicle operations into account, equipment consists of three major parts:

10.5.5 Security and Access Control • Launch vehicle stand-alone operations, duration 16 days The security requirements for Plesetsk will • Spacecraft operations, duration de- be defined in the joint launch site security pending on Customer needs, average plan. This document considers the re- duration 14 days quirements from the Russian side as well as the requirements of the Customer state • Combined operations, duration 14 days government. The spacecraft and its support equipment will arrive on day L-29 at the Russian port- of-entry, accordingly the spacecraft recep- 10.5.6 Safety tion team will usually arrive there earlier to The safety regulations (chapter 9) define prepare for the delivery and the ongoing the rules applicable to all operations and transport to the launch site. the constraints to be observed in the definition and performance of launch vehicle and spacecraft operations.

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Figure 10-11 Schedule of operations.

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The combined operations start on day L-14 usually with mating the spacecraft to the 10.5.7.2 Spacecraft Operations launch vehicle adapter. The spacecraft operations at the launch The launch vehicle processing is per- site nominally start on day L-19 with the formed in parallel with the space vehicle arrival of the spacecraft and spacecraft processing. The major launch vehicle GSE container. processing, spacecraft operations and combined launch vehicle / spacecraft tasks The spacecraft autonomous operations are are summarised in the operations schedule conducted in the spacecraft processing (Figure 10-12). area of the clean room bay of MIK. The order of spacecraft fuelling and mating

10.5.7.1 Launch Vehicle Operations in MIK and the spacecraft to the adapter may be at Launch Complex changed depending on mission specific preferences. The stand-alone launch vehicle operations at the launch site typically start on day L-28 with the arrival of the Rockot components 10.5.7.3 Combined Operations in MIK at the MIK processing facility and end on The combined operations of launch vehicle day L-14 with the start of combined opera- and spacecraft in MIK nominally start on tions. day L-14 with mating the spacecraft to the Before the components of the upper com- adapter to configure the payload stack. posite are prepared for spacecraft integra- The stack integration is usually be per- tion and assembly, an electrical Breeze- formed in the spacecraft processing area KM / booster interface check-out is per- (room 101A). All upper composite opera- formed at the launch pad. For this purpose, tions, i. e. spacecraft / Breeze-KM and the upper composite will be assembled, fairing assembly, are performed in vertical transported to the launch pad and stacked orientation. on the second stage. All operations dedi- The independent preparations of all upper cated to the electrical Breeze-KM / booster composite components are completed interface check-out activities are similar to when the payload stack is transferred from the launch operations, but do not include the spacecraft processing area to the the fuelling of the upper stage and are upper composite integration area of the performed without the spacecraft. clean room bay on a dedicated dolly. The The main launch vehicle stand-alone op- prepared and fuelled Breeze-KM is as- erations up to the start of the combined sembled on the mobile integration table operations are shown in Figure 10-12 and and the fairing is separated into halves. the sequence in Figure 10-13. The specific combined operations tasks are shown in Figure 10-12 and Figure 10-13.

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Payload Encapsulation and Upper Composite Integration

Figure 10-12 Flow of combined operations in MIK. rupted for the second time during Upper 10.5.7.4 Combined Operations at the Launch Composite installation onto the booster. Pad The third interruption takes place when the stiffening ring is removed and the fourth The upper composite with the spacecraft shut-down occurs when the transport arrives at the launch pad on day L-7 for launch container extension is installed on mating. The air conditioning of the space- the transport launch container. The tasks craft will be stopped four times for short performed at the launch pad, such as periods during ILV integration with each Rockot booster erection and mating of the shut-down period lasting no more than one upper composite are shown in Figure 1-20 hour. Air conditioning is interrupted for the to 1-23 and schematically in Figure 10-13. first time when the Upper Composite is being lifted into the service tower. During A Launch Readiness Review (LRR) on day ongoing Upper Composite preparations for L-3 the so-called “State Commission”. mating, the air conditioning system is gives permission for booster fuelling. activated. The air conditioning is inter-

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Figure 10-13 Launch pad operations.

10.5.8 Launch Day be defined in the JOP. The previous launch decision may be affected at any time dur- After the performance of all electrical ing the final count-down by modifications or check-out operations and the telemetry anomalies of the launch configuration. All readiness check carried out on the launch the modifications of the previous launch day, the State Commission decides about configuration must be reported in real time the launch and the start of the count-down. to the Operations Leader. The constitution of the management envi- sioned on the launch day, as well as their The Launch Operations Leader observes role during countdown, will be defined in the GO / NO-GO criteria and holds the the JOP. The chairman of the commission count-down, if necessary. will decide on the readiness for launch About 10 minutes before launch the ser- after readiness report of all participated vice tower is withdrawn, whereas the air parties: conditioning for the spacecraft will continue • Weather conditions until 30 seconds before launch. Mechanical access to the payload after encapsulation • Launch pad status is not planned as a standard service. How- • Launch vehicle ground tracking sta- ever, access via umbilical connectors will tions be provided during any operation phase after encapsulation, e. g. for battery trickle • Launch vehicle ground measurements charging and communication. • Spacecraft status All operations, boundary conditions, com- • Communications network munication links and other relevant infor- • Launch conditions (launch window, GO mation are prepared, agreed and docu- / NO-GO criteria, launch abort) mented in the Launch Operation Schedule (LOS), the guiding document for the launch • Launch vehicle status day. The Launch Operations Leader conducts the launch according to the criteria that will

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10.5.9 Abort Re-Cycle/ Return-to-Base 10.6 Accommodation and Leisure Operation Activities

In the case where the launch has to be Customers will be accommodated in the postponed, the Launch Operations Leader Hotel Rockot in the town of Mirny (Figure requests an agreement from the various 1-35). Mirny is the Cosmodrome's main management representatives. supporting town and has a well-developed If the count-down is held on the scheduled social infrastructure. launch day inside the launch window with- The international standard Hotel Rockot out any malfunctions of the launch vehicle was refurbished in 1999 in order to satisfy or spacecraft systems, the launch will be all needs of EUROCKOT Customers. The postponed at least by one hour required for common areas of the hotel comprise a launch systems readiness and at the most meeting room, TV lounge that can be by 48 hours, for SSO launches usually by arranged as a fitness room as well as a bar one day. During this delay, the launch and restaurant. vehicle can remain filled on the launch table, assuming the environmental tem- Each hotel room contains a bathroom, a perature is within the allowable range. desk, refrigerator, telephone, TV set able to receive Russian, local and satellite TV In the case of launch delay, thermal condi- programs and LAN outlet. The telephone tioning for the upper composite is provided. link may also be used for dial-up to a local The thermal conditioning system will be internet server and e-mail account. In total switched on one minute after launch can- there are 39 guest rooms available. cellation. One guest room on the second floor can If any malfunction is detected that could be used as a Customer office. A LAN patch not be solved in-situ, related either to the panel leading to each room and a 64 kbps launch vehicle or to the spacecraft sys- modem interface to PBX are terminated in tems, the booster will be defueled and the the entrance area of this room. The hotel upper composite will be removed from the LAN is connected to the processing area booster unit and transported back to MIK. LAN. All defects and failures on the spacecraft are repaired within the spacecraft process- For the safety of the guests, the hotel has ing area, whereas all defects and failures a fire alarm system. Each room is of Breeze-KM are repaired in the upper equipped with fire detectors and there are composite integration area of the clean smoke detectors on the corridors. In the room bay. After repair and check-out are event of an alarm, an audible alert will be finished, the upper composite will be inte- sounded in the lobby area and on each grated and transported back to the launch floor. Fire hoses, plumbing and emergency pad, re-integrated on top of the booster exits are installed on each floor. unit, and the launch preparation cycle will In the immediate vicinity of the hotel there begin again. The booster unit remains at are two stadiums, tennis, volleyball and the pad during this process. basketball courts, and an indoor athletics

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complex accommodating a swimming pool, a gym and a fitness centre. Trips to wildlife areas or historic places, sightseeing, sports events and games, jogging and use of the athletic complex facilities and sauna can also be arranged.

10.7 Medical Care

A well-equipped military hospital can treat up to 200 patients. The medical team is trained to the highest standard available in Russia. Ambulances are available. Com- panies offering rapid medical evacuation services to Western Europe can be arranged upon request.

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Chapter 11 Baikonur Cosmodrome

Table of Contents

11 Baikonur Cosmodrome...... 11-1

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11 Baikonur Cosmodrome

The modification of KSRC dedicated facilities at Baikonur Cosmodrome for EUROCKOT launches is possible but currently not being pursued. Although Rockot launches have occurred in the past from Baikonur Cosmodrome, it is currently no longer operational for Rockot launches. Activation of the site for Rockot/Breeze-KM requires upgrades and modifications to existing facilities. A decision for site activation will be made on a case-by-case basis should the need arise.

Generally, the extent and quality of the EUROCKOT facilities and services at Baikonur will at least be similar to those in the Plesetsk Cosmodrome, and in some respects even better.

Baikonur is particularly suited for serving inclinations in the 50° range, since these cannot be efficiently reached from Plesetsk due to its northerly latitude.

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Chapter 12 Items to be Delivered by the Customer

Table of Contents

12 Items to be Delivered by the Customer ...... 12-1

12.1 General Documents ...... 12-2 12.1.1 Interface Requirements Document...... 12-2 12.1.2 Orbit Tracking Operation Report ...... 12-3 12.2 Input to Mission Design and Mission Analysis ...... 12-3 12.2.1 Response to Questionnaire...... 12-3 12.2.2 Spacecraft Dynamic Model ...... 12-5 12.2.3 Thermal Model ...... 12-5 12.3 Safety ...... 12-6 12.4 Payload Environmental Test Documents ...... 12-6 12.5 Operations Documents for Spacecraft ...... 12-6 12.6 Contractual / Higher Level Documents...... 12-7 12.7 Models, GSE ...... 12-7 12.8 Hardware, Software and Document Time Schedule ...... 12-8

List of Figures

Figure 12-1 Interface Requirements Document table of contents...... 12-2 Figure 12-2 Spacecraft Operations Plan table of contents...... 12-7 Figure 12-3 Dates of the Customer's documents, software and hardware supply...... 12-9

List of Tables

Table 12-1 Summary of documents to be supplied by the Customer...... 12-1

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12 Items to be mission insignias for the launch vehicle or container-external surfaces, are to be Delivered by the agreed on in the first two mission prepara- Customer tion phases. Wherever possible, submission of data in electronic format is pre- ferred by EUROCKOT in order to improve This chapter summarises and describes shipment times and accessibility of the the main software (documents, data and data. software-models) and hardware (flight units, dummies and ground equipment) Table 12-1 provides a summary of all which shall be supplied by the Customer documents to be supplied by the Cus- as a minimum. The times and destinations tomer during the various mission phases. for shipments are also described. Delivery Further explanations regarding their defini- times and destinations of additional items tion can be found in the following sections. which the Customer wants to use, e. g.

Date (typically) Documents to be provided Preliminary Final Interface Requirements Document (IRD) L - 24 months I: L-18 months Safety submission (Phase I, II, III) III: L- 6 months II: L-12 months Spacecraft mechanical environment test plan L - 16 months Spacecraft dynamic model (Preliminary) L - 23 months L - 11 months Spacecraft thermal model (Preliminary) L - 23 months L - 11 months Response to questionnaire: Input to Mission Design and Mission L - 23 months L - 11 months Analysis Launch license documentation L - 10 months Spacecraft Operations Plan L - 17 months Spacecraft mechanical environment qualification test results L - 6 months Technical readiness documentation: SC technical readiness certificate, SC readiness package for launch campaign opera- L - 6 months L - 3 months tions within the integration facility, launch pad and launch Final customs documentation SC shipment - 1 month Final spacecraft mass properties L - 7 days Orbital Tracking Operation Report L + 2 weeks

Table 12-1 Summary of documents to be supplied by the Customer.

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12.1 General Documents

1 INTRODUCTION 12.1.1 Interface Requirements Document 1.1 Objective 1.2 Mission Description Interface Requirements Document L-24 months 1.3 Project Summary

The Interface Requirements Document 2 APPLICABLE DOCUMENTS (IRD) will be the technical baseline docu- 2.1 Government Documents ment for the first mission phase as long as 2.2 Customer Documents no Interface Control Document has been 2.3 Reference Documents established and agreed. The IRD will be created by the Customer and is generally 3 REQUIREMENTS one of the parts of the technical annexes of 3.1 MISSION REQUIREMENTS 3.1.2 Launch Requirements the launch services contract. EUROCKOT 3.1.2.1 Launch Date can supply a generic IRD template for cus- 3.1.2.2 Launch Time 3.1.3 Injection Orbit Requirement tomers to use, if required. 3.2 SEPARATION REQUIREMENTS With this document, the Customer will also 3.2.1 Separation velocity 3.2.2 Angular velocities describe the mission and spacecraft char- 3.2.3 Separation monitoring acteristics as already defined. This focuses 3.3 INTERFACE REQUIREMENTS on mass properties, interface dimensions, 3.3.1 Mass Properties 3.3.2 Mechanical Interfaces and mission and orbit characteristics in 3.3.2.1 Static Envelope and particular. Within the outline provided by Clearances EUROCKOT, all chapters which have to be 3.3.2.2 Satellite to Launch Vehicle Interface completed with information for contract 3.3.2.3 Satellite Stiffness signature will be marked. 3.3.3 Electrical Interfaces 3.3.4 RF Link (if applicable) Figure 12-1 shows the typical table of con- 3.4 LAUNCH SITE OPERATIONAL REQUIREMENTS tents for this. Nevertheless, modifications 3.4.1 Transportation according to dedicated demands of the 3.4.2 Payload Processing Facility mission can be implemented on the basis 3.4.3 Launch Pad of joint agreements. 3.5 MECHANICAL ENVIRONMENTS 3.5.1 Static Loads 3.5.2 Low Frequency Vibration 3.5.3 Acoustic Noise 3.5.4 Separation Shock

APPENDICES AND DRAWINGS

Figure 12-1 Interface Requirements Document table of contents.

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− Requirements (if any) of spacecraft 12.1.2 Orbit Tracking Operation Report attitude to the sun during coast phases of upper stage

Orbit Tracking Operation Report L+2 weeks − Manoeuvres during upper stage coast phase, e.g. thermal manoeu- In order to confirm Rockot performance vres. with regard to orbit injection accuracy, the − Separation attitude of spacecraft Customer is requested to submit space- − Requirements to the launch vehicle craft tracking data after third stage burnout after separation, e.g. collision before and subsequent to separation as far avoidance manoeuvres, constraints as such data are available. This must in- on thrusters operations etc. clude a complete set of orbital parameters − Requirements to the launch window and their estimation accuracy. and allowable launch window duration (if not specified by other parameters) 12.2 Input to Mission Design and • Spacecraft characteristics Mission Analysis − Payload designation − Dimensions of spacecraft stowed in 12.2.1 Response to Questionnaire launch configuration and when deployed Response to Questionnaire: − Mass and inertial characteristics of Input to Mission Design and Mission Analysis Preliminary: L-23 months dry and fuelled spacecraft Final: L-11 months − Propellant characteristic, viscosity, density etc. In addition to the dynamic and thermal models for coupled loads and thermal − Fuel sloshing analysis inputs (if applicable) analyses, EUROCKOT requires additional input data and information to adequately − Thermal model perform the preliminary and final mission − Dynamic model design and analyses. It should be noted • Ground and launch environments that some of this data will probably be contained within the customer supplied IRD − Quasi-static and dynamic loads in and the resulting ICD that is established. flight However, all the required data is repeated − Transportation loads here for completeness. − Separation shock loads • Flight programme specification includ- − Acoustic loads ing − Ground temperature / humidity con- − Required injection orbit and allow- straints able errors − Flight temperature constraints / fairing internal surface temperature constraints

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− Spacecraft cleanliness require- − Spacecraft transportation provisions ments, i. e. particles, surface − Processing area requirements cleanliness, organics (if applicable) − Spacecraft fuelling area require- − Pressure / venting constraints ments within payload fairing − Personnel accommodation office in − Free molecular heating rate con- facilities / hotel requirements straints after payload fairing jettison. − Technical support requirements at launch base − RF/ EMI constraints − Communication requirements, e.g. Spacecraft Interfaces • LAN interfaces, internet access, − Spacecraft coordinate system and mobile walkie-talkies etc. reference to launcher coordinate − Spacecraft ground support equip- system ment quantity / size etc. − Spacecraft static envelope (stowed, − Required consumables (gases etc.) in launch configuration), critical points of spacecraft envelope rela- − Security requirements tive to fairing. • Campaign schedule/ operations − Preferences for spacecraft location and clocking within payload fairing • Drawings of spacecraft handling units and transport containers − Flight mechanical interfaces, including spacecraft to launcher interface • Requirements for installation flanges / points. − Ground mechanical interfaces, e. g. • Points for hoisting and fixing to handling dolly, fuelling platform • Requirements for the separation sys- etc. tem − Electrical interfaces including quantity, type and location of umbilical • Data on the payload elements which connectors have to be jettisoned or deployed − Umbilical connector separation force • Pyrotechnic devices and related constraints − Content and parameters of umbilical lines for flight • Orbital parameters for the payload − Electrical interfaces for ground op- • Requirements for injection accuracy erations and payload orientation prior to its de- − Telemetry parameters to be re- ployment corded during flight − Grounding and bonding require- • Acceptable range for thermal environments ments during the payload injection phase − Injection orbit data reporting formats • Requirements for protection of optical • Launch site requirements surfaces

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• Thermal control requirements Other presentations of the mathematical models, e. g. a spring mass model, are to • Parameters of payload / ground sup- be agreed with EUROCKOT. port equipment interfaces

• Characteristics of the payload tele- 12.2.3 Thermal Model metry and telecommand system and other RF systems Thermal Model Preliminary L-23 months • Payload and related GSE input data: Final L-11 months

− Allowable thermal conditioning in- Section 8.5 describes which thermal envi- terruptions for the payload, batteries ronment studies are required to verify and propellant containers thermal compatibility through out the mis- − Payload processing cycle duration sion. This study will be implemented using in integration facility and at launch a thermal model provided by the Customer. pad As this study covers the period from inte- − Payload ambient temperature, hu- gration of the payload onto the dispenser midity and contamination control within the integration facility, up to space- requirements during operations craft separation, the Customer has to pro- − Spacecraft battery charging/trickle vide the following: charging cycles in integration facility and at launch pad. • A thermal model of the spacecraft containing

12.2.2 Spacecraft Dynamic Model − a description of the thermal nodes (heat capacities, mass type, etc.)

Spacecraft Dynamic Model − internal thermal couplings of nodes Preliminary L-23 months (conductive, radiative and convec- Final L-11 months tive) As described in section 8.4, structural − heat dissipation for all applicable compatibility will be demonstrated with modes of operation during the covered mission phases preliminary and final CLA. Customer inputs, in particular structural models of the • interface descriptions (areas of contact, spacecraft, are requested for both prelimi- conductive and/or radiative properties) nary and final CLA steps. • thermal requirements for the environment to be fulfilled during integration, The spacecraft mathematical models must launch and flight be provided by the Customer in the form of For detailed descriptions, refer to section stiffness matrices and masses of non-fixed 8.5 and EUROCKOT document ESPE- structures, mathematically reduced to a 0009. Craig-Bampton model. For detailed descriptions, refer to section 8.4.1 and to EUROCKOT document ESPE-0008.

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12.3 Safety setsk Cosmodrome and for launch on the Rockot/Breeze-KM launch vehicle. Safety Submissions Step 1 L-18 months Step 2 L-12 months The spacecraft readiness data package Step 3 L-6 months providing information necessary to justify During the mission phases, safety submis- spacecraft readiness for ground operations sions have to be provided by the Customer and flight certificate. in three steps. The content and format of the data to be supplied are described in more detail in chapter 9 and in the 12.4 Payload Environmental Test Documents EUROCKOT Safety Handbook EHB-0004.

Generally, all areas generating risks for Spacecraft Mechanical Environment personal safety such as pressurised sys- Qualification Test Report L-6 months tems, explosives (propellants, pyrotechni- After the performing of structural qualifica- cal devices, etc.), radioactive material, RF tion, test results shall be submitted to sources and toxic substances have to be EUROCKOT for a review of compliance covered, as well as safety-relevant opera- with the structural model supplied for the tions to be performed during ground prepa- coupled loads analysis. If any discrepan- rations. It has to be proven with all avail- cies regarding loads, strength or stiffness able information, how risks to people in- were identified during qualification testing, volved can be minimised to acceptable corrective actions have to be agreed on. levels, which safety factors have been applied and how they have been or will be Certainly, just as for the acceptance test verified, and which precautions are envis- results below, the extent to be provided is aged. subject to mutual agreements as far as proprietary or technology export issues are SC Technical Readiness Documentation Preliminary L-6 months involved. Final L-3 months

Spacecraft Technical Readiness Docu- 12.5 Operations Documents for mentation submitted by the Customer Spacecraft should include: For the organisation of work within the • The spacecraft technical readiness integration facility and on the launch pad, certificate for launch campaign opera- the following documents are required: tions within the integration facility, on

the launch pad and for launch. This Spacecraft Operations Plan L-17 months certificate ensures that the spacecraft is designed and checked in compliance The purpose of the SOP is to define the with the ability to take all environmental activities to be performed on the spacecraft loads, specified in ICD, and is ready for during the launch campaign and the rele- launch campaign operations at Ple- vant support and facilities required at the range. The document shall also be for-

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warded to EUROCKOT to establish the • An End User Certificate that briefly launch campaign Joint Operations Plan describes the intended purpose of the (JOP) which is agreed between cos- mission, end user and the instrumenta- modrome authorities, EUROCKOT, KSRC tion of the spacecraft is to be provided and the Customer. The JOP is further dis- 10 months prior the launch. cussed in chapter 7. An outline version of a • The Customer is also responsible for typical SOP document is given in Figure obtaining the export license and the 12-2. appropriate approval for use of radio For spacecraft shipment and customs frequencies in the intended orbit in a clearance the customer has to prepare and timely manner. deliver the final pro-forma invoice accompanied by the detailed packing list of the spacecraft shipment not later than one 12.7 Models, GSE month prior to the spacecraft shipment. As a minimum, two hardware models have Information about hazardous materials in to be available for testing (for more details internationally accepted formats also have refer to section 7.2.1.2. to be provided, if applicable.

1. Introduction Mass Frequency Simulator L-12 months 2. Applicable Documents 3. General A spacecraft model simulating at least me- 4. Operations/ Baseline Schedule including • Test plan, day-by-day planning chanical interfaces, mass and CoG posi- • Preparations and check-out to be carried out tion has to be provided by the Customer. It in the integration facility also has to be mutually agreed how far and • Assembly of the payload with the upper stage • Payload fuelling procedure SC with which tolerances, MoIs and stiffness • Payload control and monitoring on the launch characteristics have to be simulated. As a pad • Warning regarding handling baseline, the main natural frequencies • Launch constraints should be simulated. • Launch window • Equipment associated with spacecraft • Electrical wiring requirements Fit Check/Dummy (TBC) L-12 months • Installations (buildings etc.) • Logistics In case of very low geometrical clearance between the spacecraft and the payload Figure 12-2 Spacecraft Operations Plan table of contents. fairing, an advanced mock-up model similar to the flight unit regarding geometrical and mechanical interfaces has to be supplied for testing at KSRC's premises in 12.6 Contractual / Higher Level Moscow. The fit of overall dimensions with Documents the accommodation envelope (upper stage cover, dispenser, adapter and fairing) As well as technical documents other in- would be verified during this test. If it is puts will be requested from the Customer obvious that there are no clearance issues for obtaining the launch license. regarding the fairing, a fairing fit check and

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a spacecraft geometrical model are not Others required. Potentially, the fit check dummy could be provided in the form of the mass Shipments of other items which the Cus- frequency simulator and geometric adapter tomer requires for ground operations (e. g. mentioned above. unit testers for integration facilities and launch pad, pathfinder spacecraft or con-

Flight Unit for Matchmate Test L-5 months tainers, special transportation and handling equipment, fuelling equipment as well as The Matchmate test with the launch vehicle personal safety equipment and fuel itself) adapter flight unit and the spacecraft flight as well as their storage and application are unit shall be performed at the facilities of matters for dedicated arrangements be- the spacecraft manufacturer in order to tween the Customer, EUROCKOT and the prove mechanical, electrical and opera- range operation organisations. tional compliance of this interface. A time slot as well as personal and technical resources have to be provided by the Cus- 12.8 Hardware, Software and tomer and/or spacecraft manufacturer of Document Time Schedule the flight unit. A summary of all hardware, software and documents to be provided by the Customer Master Gauge / Drill Templates L-18 months is shown in Figure 12-3. For point attachment interfaces it has to be Generally, EUROCKOT is open to agree- mutually agreed whether and by whom ments on any modification imposed by tools will be provided to enable precision special mission requirements if it is possi- positioning of attachment/fixing points at ble to consider it within the overall sched- the spacecraft and the dispenser. This will ule. not be necessary, if the same degree of precision can be achieved by fulfilling drawing requirements only or if a clamp band separation system is used.

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Figure 12-3 Dates of the Customer's documents, software and hardware supply.

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