Land Launch User s Guide

Transcription

1 Land Launch User s Guide Original Release Date: 28 July 2004 Initial Revision The Land Launch User s Guide has been cleared for public release by the Department of Defense, Directorate for Freedom of Information and Security Review, as stated in letter 04-S-1323, dated 28 July Prepared by Sea Launch for distribution by: Boeing Launch Services One World Trade Center, Suite 950 Long Beach, CA 90831, USA on behalf of the Sea Launch Company, L.L.C. Copyright pending by Sea Launch Company, L.L.C.

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3 Land Launch User s Guide REVISIONS LTR DESCRIPTION DATE APPROVAL NC Initial release July 2004 James Ellinthorpe Program Manager iii Initial Release

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5 Land Launch User s Guide TABLE OF CONTENTS 1. INTRODUCTION Purpose Page 1.1 Overview of the Land Launch System What the System Includes Advantages to the Customer Timeline Baikonur Cosmodrome Land Launch Organization Overview Sea Launch Company, LLC The Boeing Company Space International Services, Ltd SDO Yuzhnoye PO Yuzhmash Design Bureau of Transport Machinery (KBTM) Center for Ground Space Infrastructure Operations (TsENKI) RSC Energia NPO Lavochkin Russian Space Agency VEHICLE DESCRIPTION Overview Design Zenit Flight History Block DM Flight History Flight Success Ratios Land Launch Zenit Design Heritage Changes Made for Sea Launch Avionics Overall Specifications and Configurations Zenit Stage Overall Configuration RD-171M Engine v Initial Release

6 Land Launch User s Guide Zenit Stage Overall Configuration RD-120 Main Engine RD-8 Vernier Engine Block DM-SLB Upper Stage Overall Configuration D58M Main Engine Attitude Control/Ullage Engines Avionics Changes Made for Sea Launch Block DM-SLB Versus the Block DM-SL Zenit-3SLB Ascent Unit Components and Integration Payload Fairing Fairing Access Characteristics Conditioned Air Supply to the Fairing Fairing Thermal Protection Payload Structure Support The Zenit-2SLB Payload Unit Components and Integration Payload Fairing Fairing Access Characteristics Conditioned Air Supply to the Fairing Fairing Thermal Protection Intersection Bay Spacecraft Adapters Unique Interfaces and Multi-Spacecraft Launches PERFORMANCE Overview Performance Ground Rules Launch Window Availability Launch Site and Accessible Orbits Site Location Accessible Orbits vi Initial Release

7 Land Launch User s Guide 3.2 Ascent Trajectory Generic Zenit-3SLB GTO Mission Mission Profile Stage 1 Flight Stage 2 Flight Block DM-SLB Powered Flight Flight Timeline Ground Track Ascent Trajectory Generic Zenit-2SLB Mission to 51.6 LEO Stage 1 Flight Stage 2 Flight Flight Timeline Flight Profile Ground Track Payload Capability Three Stage Zenit-3SLB Geosynchronous Transfer Orbit MEO, HEO, Circular, and Elliptical Orbits High-Energy and Earth-Escape Trajectories Payload Capability Two Stage Zenit-2SLB Circular LEO Orbits Elliptical Orbits Coast Phase Attitude Maneuvers Zenit-3SLB Zenit-2SLB Injection Accuracy Spacecraft Separation and Post-Separation Events Zenit-3SLB Separation Event Separation Capabilities CCAM State Vector Delivery Zenit-2SLB Separation Event Separation Capabilities CCAM for Second Stage State Vector Delivery vii Initial Release

8 Land Launch User s Guide 4. SPACECRAFT ENVIRONMENTS Overview Ground and Flight Environments Reference Coordinate System Environmental Monitoring Structural Loads Overview Quasi-Static Load Factors, Ground Handling, and Transportation Quasi-Static Load Factors, Flight Sinusoidal Equivalent Vibration During Flight Random Vibration Ground Random Vibration for Components Near Spacecraft Interface Flight Random Vibration Environment Acoustics Fairing Space Average Sound Pressure Levels Shock Overview Zenit-3SLB Zenit-2SLB Electromagnetic Environment Overview Coordination Ambient Cosmodrome Electromagnetic Environment Launch Vehicle Radio Equipment Radio Frequency Environment at the SC Separation Plane Spacecraft Thermal and Humidity Environments Introduction Ground Thermal and Humidity Environments General Overview, Ground Thermal and Humidity Environments Facility Clean Air Systems Transportation Clean Air Systems Launch Pad Clean Air System Impingement Velocity of Airflow Upon SC Surface Flight Thermal Environments General Overview, Flight Thermal Environment viii Initial Release

9 Land Launch User s Guide 4.7 Pressure Venting Overview Pressure Decay Rate Pressure Differential at Fairing Jettison Contamination Contamination Control During Ground Processing Contamination Control During Flight Fairing Design Features to Minimize Contamination Plume Impingement SPACECRAFT INTERFACES Mechanical Interfaces Mass Properties and Modal Frequencies Spacecraft Mass and Longitudinal Center of Gravity Location Spacecraft Center of Gravity Radial Offset Modal Frequencies Payload Fairing Mechanical Interfaces Payload Fairings Useable Volume Useable Volume Inside Payload Structure Access Doors RF Windows Customer Insignia Spacecraft Adapters Saab Spacecraft Adapters Zenit-2 Adapter for Use with Zenit-2SLB Multi-Satellite Dispensers for Use with Zenit-2SLB Electrical Interfaces Overview Hard Line Links (Spacecraft Umbilical) Umbilical Circuits Umbilical Use During Processing and Launch Umbilical Connectors Radio Frequency Links In-Flight Commands, Measurements and Telemetry General Separation Verification Satellite Environments Measurements Commands Electrical Power for EGSE Ground Power ix Initial Release

10 Land Launch User s Guide Uninterruptible Back-up Power Bonding and Grounding Bonding Grounding LAND LAUNCH FACILITIES Overview Transportation of Personnel and Cargo to and From Baikonur Krainy Airport Yubileiny Airport Transportation at the Cosmodrome Site 31 Payload Processing Facility Overview Buildings 40/40D, PPF Building 40D Office Areas Building 44, HPF Site 254 Payload Processing Facility Overview Site 254 PPF Layout Site 254 PPF Features Zenit Technical Complex Site Overview Integration Area Layout/Features Spacecraft Equipment Room Customer Office Areas Zenit Launch Complex (LC) Site Overview Launch Complex Automated Systems Customer EGSE Room (Bunker) Command Center Cosmodrome Amenities Visa and Access Authorization Customs Clearances Transportation Consumables Security Schedules External Communications Medical Care Accommodations and Dining x Initial Release

11 Land Launch User s Guide 7. SPACECRAFT DESIGN & VERIFICATION REQUIREMENTS Overview Additional Spacecraft Design Consideration Constraints on Spacecraft Transmitting and Receiving Horizontal Handling Safety Design Considerations Pressurized Systems Ordnance Systems Ground Support Equipment (GSE) Considerations Spacecraft Structural Capability Flexibility Spacecraft Structural Capability Factors of Safety Test Verified Model Required for Final CLA Test Requirements Modal Survey Test Static Loads Test Sine Vibration Testing Acoustic Testing Shock Qualification Matchmate Test MISSION INTEGRATION AND OPERATIONS Overview Mission Management Mission Manager Mission Team Roles and Responsibilities Mission Documentation and Schedule Overview Integration Documentation Spacecraft/Land Launch System Interface Control Document ICD Verification Matrix Mission Integration Schedule Mission Analyses Mission Analyses Operations Planning Launch Campaign Planning xi Initial Release

12 Land Launch User s Guide 8.5 Launch Campaign Overview Spacecraft Arrival and Transport Spacecraft Processing Spacecraft Fueling Spacecraft Mating with Launch Vehicle Elements in PPF Integration Bay Launch Vehicle Autonomous Processing Mating with Zenit Stages Integrated Testing Transfer Readiness Review and Transfer to Launch Pad Launch Pad Operations Launch Readiness Review Propellant Loading Second Go Poll and Launch Launch Control Center Post-Flight Activities Safety Quality Assurance General Hardware Review APPENDIX A USER QUESTIONNAIRE... A-1 Spacecraft Physical Characteristics... A-2 Spacecraft Orbit Parameters... A-3 Guidance Parameters... A-3 Electrical Interface... A-4 Thermal Environment... A-8 Dynamic Environment... A-9 Ground Processing Requirements... A-10 Contamination Control Requirements... A-14 APPENDIX B SEA LAUNCH STANDARD OFFERINGS AND OPTIONS...B-1 Standard-Offering Hardware...B-1 Launch Vehicle...B-1 Payload Fairing and Spacecraft Adapter...B-1 Electrical interfaces...b-2 xii Initial Release

13 Land Launch User s Guide Standard-Offering Launch Vehicle Performance...B-2 Orbit and Mass...B-2 Orbit Accuracy...B-2 Standard-Offering Launch Services...B-2 Mission Management...B-2 Meetings and Reviews...B-3 Documentation...B-3 Mission Integration...B-4 Interface Test...B-4 Standard-Offering Facilities And Support Services...B-5 Payload Processing Facilities...B-5 PPF Communication...B-6 PPF Security...B-6 PPF Support Services...B-6 Launch Vehicle Integration Facility, Area 42...B-7 Launch Complex (LC) Facilities, Area 45...B-7 Launch Complex Communications...B-7 Launch Complex Area 45, Security...B-7 Environmental Controls...B-8 Range Services...B-8 Logistics Support...B-8 Optional Services...B-8 Mission Analysis...B-8 Interface Tests...B-9 Support Services...B-9 Facilities...B-9 Materials...B-9 xiii Initial Release

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15 Land Launch User s Guide LIST OF FIGURES Page 1-1 Sea Launch Zenit-3SL Mission Integration Timelines The Baikonur Cosmodrome Land Launch Organizational Structure The Zenit Vehicle(Yuzhnoye) Zenit Complex at Baikonur (KBTM) The Block DM (Energia) The Zenit-3SLB The Zenit-2SLB Cosmonaut Access Tower at the Zenit Launch Complex Zenit Stage 1 and Stage 2 Configuration Land Launch Zenit Stage RD-171M Engine Zenit Second Stage Block DM-SLB Zenit-3SLB Ascent Unit Zenit-2SLB Payload Unit Zenit 2 Spacecraft Adapter Flight Corridors for Land Launch from Baikonur Three-burn Block DM Mission Profile to GTO Approved Land Launch Ground Track and Drop Zones for GTO Injection Ground Track for Generic Zenit-3SLB GTO Mission Typical Ascent Profile to ISS Orbit Flight Ground Track for Zenit-2SLB Mission to ISS Orbit Zenit-3SLB Payload Capability to GTO Zenit-3SLB Performance to Circular Orbits Zenit-3SLB Performance to Elliptical Orbits Zenit-3SLB High-Energy and Earth Escape Payload Capability Zenit-2SLB Payload Capability for Circular Low Earth Orbits Zenit-2SLB Performance to Elliptical Orbits Reference Coordinate System to Define Spacecraft Environments Typical Quasi-Static (Max Expected) Design Loads in Flight Random Vibration Environment During Flight Max Expected Acoustic Pressure Envelope inside Zenit-2SLB Fairing Max Expected Acoustic Pressure Envelope inside Zenit-3SLB Fairing a Zenit-3SLB Spacecraft SRS with Standard SAAB 937-mm Adapter b Zenit-3SLB Spacecraft SRS with Standard SAAB 1194-mm Adapter c Zenit-3SLB Spacecraft SRS with Standard SAAB 1666-mm Adapter d Zenit-2SLB Spacecraft SRS with Standard SAAB 2624-mm Adapter a Ambient Electromagnetic Environment within PPF Site b Ambient Electromagnetic Environment within PPF Area Ambient Electromagnetic Environment within ILV Assembly Bldg Ambient Electromagnetic Environment at Zenit Launch Complex xv Initial Release

16 Land Launch User s Guide 4-10 Max Field Intensity Levels at SC Separation Plane, Zenit-2SLB Max Field Intensity Levels at SC Separation Plane, Zenit-3SLB Zenit-2SLB Ascent Unit Air-Conditioning and Venting Scheme Zenit-3SLB Ascent Unit Air-Conditioning and Venting Scheme Zenit-3SLB Free Molecular Heating Environment Zenit-2SLB Ascent Venting Scheme Zenit-3SLB Ascent Venting Scheme Typical Zenit-2SLB Fairing Internal Pressure Profile During Ascent Typical Zenit-3SLB Fairing Internal Pressure Profile During Ascent Typical Zenit-2SLB Fairing Internal Pressure Profile Decay Profile Typical Zenit-3SLB Fairing Internal Pressure Profile Decay Profile Fairing for Zenit-3SLB Fairing for Zenit-2SLB General Lay-out of the 4.1-meter 17S72 Fairing (Zenit-3SLB) General Lay-out of the 3.9-meter Fairing (Zenit-2SLB) Spacecraft Static Envelope within Zenit-2SLB Fairing Spacecraft Static Envelope within Zenit-3SLB Fairing Locations for Access Doors, Zenit-3SLB Payload Fairing Locations for Access Doors, Zenit-2SLB Payload Fairing Location of Land Launch Facilities at Baikonur Krainy Airport Spacecraft Off-Load at Yubileiny Airport Ascent Unit Transportation with Thermostating Car Area 31 Partial Facility Lay-out Lay-Out of Buildings 40 and 40D SC Processing and Joint Operations Area in Buildings 40 and 40D Hazardous Processing Facility, Building 44, at Site Lay-out of SC PPF at Site 254 with Proposed Adjacent Building Encapsulation Operations in Site 254 Room North Rail at the Zenit Technical Complex, Site Clean Room at Area Lay-out of the Zenit Launch Complex, Area Location of Room 114 (Customer EGSE Room) Customer Location Options in the Launch Command Center Hotel 1 at Site 2Zh Near the Site 254 PPF Maximum Intentional Spacecraft Electric Field Impingement on Launch Vehicle Electrical and Mechanical Matchmate Test xvi Initial Release

17 Land Launch User s Guide LIST OF TABLES Page 2-1 Sea Launch Stages, Cumulative Flight History, All Related Configurations Reliability and Flight History, Sea Launch Configuration Only Land Launch Zenit Specifications Block DM-SLB Specifications Launch Operational Features Zenit Launch Azimuths and Inclinations from Baikonur Accessible Orbits on Land Launch Flight Timeline GTO by Zenit-3SLB Flight Timeline Zenit-2SLB ISS Mission Zenit-3SLB Payload Capability to GTO Zenit-3SLB Performance to Circular Orbits Zenit-3SLB Performance to Elliptical Orbits Zenit-3SLB High-Energy and Earth Escape Payload Capability Zenit-2SLB Payload Capability for Circular Low Earth Orbits Zenit-2SLB Performance to Elliptical Orbits Zenit-2SLB and Zenit-3SLB Orbital Insertion Accuracy Zenit-3SLB Direct GEO Insertion Accuracy Spacecraft Motion after Separation Single Payload, Zenit-2SLB Spacecraft Motion after Separation Multiple Payloads, Zenit-2SLB Maximum Quasi-Static Accelerations During Ground Operations Sinusoidal Vibrations at Spacecraft Interface Random Vibration during Ground Transport, not in Spacecraft Container Random Vibration Environment During Flight Maximum Expected Acoustic Pressure Envelope Inside Fairings Zenit-3SLB Spacecraft SRS with Standard SAAB Adapters Zenit-2SLB Spacecraft Shock Response Spectra (SRS) Characteristics of the Sirius Transmitters (Zenit-2SLB and Zenit-3SLB) Characteristics of BITC-B Telemetry Equipment (Zenit-3SLB Only) Characteristics of Glonass Receiver (Zenit-2SLB and Zenit-3SLB) Max Field Intensity Levels Generated by Launch Vehicle at SC Separation Plane, Without Fairing Attenuation Spacecraft Ground Thermal and Humidity Environment Flight Thermal Environments Fairing Internal Surface Cleanliness Levels at Encapsulation Expected Spacecraft Mass and CG Limits Zenit-3SLB Expected Spacecraft Mass and CG Limits Zenit-2SLB Recommended Spacecraft Fundamental Frequencies Standard SAAB Ericsson Space Spacecraft Adapters Spacecraft Umbilical Links Umbilical Hook-up Locations and Availability Characteristics of Commands from Launch Vehicle to Spacecraft Electrical Power Supplies for Customer EGSE Uninterruptible Power Supply for Customer EGSE xvii Initial Release

18 Land Launch User s Guide 7-1 Factors of Safety and Test Options Sine Vibration Amplitudes and Sweep Rates Spacecraft Acoustic Margins and Test Durations Typical Mission Integration Schedule Typical Launch Campaign Schedule xviii Initial Release

19 Land Launch User s Guide Abbreviations and Acronyms A A 0 A/C ASTM ATB AU B BER BLS BPS C C 3 CA CC CCAM CCTV CDR CG CIS CLA CM CVCM DB DC DP EGSE EMC ESD F FM FMH FSA FT G GEO ampere(s) azimuth air conditioning American Society for Testing and Manufacture Assembly & Test Building ascent unit Baikonur bit error rate Boeing Launch Services bits per second Celsius or Centigrade velocity squared at infinity California command center contamination and collision avoidance maneuver closed circuit television critical design review center-of-gravity Commonwealth of Independent States coupled loads analysis centimeter(s) collected volatile condensable material decibel(s) direct current dew point electrical ground support equipment electromagnetic compatibility electro-static discharge Fahrenheit frequency modulation free molecular heating Federal Space Agency foot/feet gravity geosynchronous or geostationary Earth orbit xix Initial Release

20 Land Launch User s Guide Abbreviations and Acronyms GOWG ground operations working group GTO geosynchronous transfer orbit GSE ground support equipment H; HR hour HEO high Earth orbit HPF hazardous processing facility HZ hertz I inclination moment of inertia ICAO International Civil Aviation Organization ICD interface control document I/F interface ILV integrated launch vehicle IRD interface requirements document ISS International Space Station K thousand(s) KBTM Design Bureau of Transport Machinery KG kilogram(s) KGF Kilogram(s) force KM kilometer(s) KN kilonewton(s) KVA kilo volt-ampere(s) KW kilowatt(s) LB(S) pound(s) LBF pound(s) force LEO low Earth orbit LC launch complex LLC limited liability company LOX liquid oxygen LRR launch readiness review LSA launch services agreement LV launch vehicle M meter(s) MS millisecond(s) ME main engine MEO medium Earth orbit MHZ megahertz xx Initial Release

21 Land Launch User s Guide Abbreviations and Acronyms MM N/A PA PCM PDR PLF PLU PPF PPM PSI PSM PSS R&D RF RH S S/C; SC SCA SCAPE SIS SOW SPST SRS T TBD TBR TM TML TRR TsENKI UPS USA USSR V VAC millimeter(s) not applicable pascal(s) pulse control modulation preliminary design review payload fairing payload unit payload processing facility parts per million pounds per square inch payload systems mass payload support structure research and development radio frequency relative humidity second(s) spacecraft spacecraft adapter self-contained atmospheric protection ensemble Space International Services, Ltd statement of work solid propellant separation thrusters shock response spectra Time tonne(s) to be determined to be revised or reviewed telemetry total mass loss transfer readiness review Center for Ground Space Infrastructure Operations uninterruptible power supply United States of America Union of Soviet Socialist Republics volt(s) velocity volt(s) of alternating current xxi Initial Release

22 Land Launch User s Guide Abbreviations and Acronyms W watt(s) 0 degrees σ sigma (one standard deviation) µ micro (10 6 ) xxii Initial Release

23 Land Launch User s Guide Section 1 1. INTRODUCTION Purpose The purpose of the Land Launch User s Guide is to familiarize members of the customer community with the Land Launch system and associated services. This document is the starting point for understanding the Land Launch spacecraft integration process, Land Launch interfaces and the overall capabilities of the system. Land Launch services are provided by the Sea Launch Company, LLC, acting in cooperation with Space International Services, Ltd (SIS), and are marketed by Boeing Launch Services (BLS). Further information may be obtained by contacting BLS directly ( 1.1 Overview of the Land Launch System Overview As its name implies, Land Launch is Sea Launch on land: the proven hardware, processes and people of Sea Launch shifted to a land-based launch site at the Baikonur Cosmodrome. Land Launch and Sea Launch complement each other by addressing different classes of payloads. Whereas Sea Launch is a heavy-lift launch system that delivers more than 6,000 kilograms to a geosynchronous transfer orbit (GTO) requiring less than 1500 meters/second to geostationary orbit (GEO), Land Launch is a mediumlift launch system that delivers 3,600 kg to an equivalent GTO. The difference in GTO performance is due to the change in launch site. There are two Land Launch configurations: The Zenit-3SLB ( B for Baikonur), a three-stage integrated launch vehicle (ILV) closely derived from the Sea Launch Zenit-3SL (Figure 1-1), is suited for delivering payloads to medium and high, circular and elliptical Earth orbits, including GTO and GEO, as well as escape trajectories. The Zenit-2SLB, a two-stage ILV based on the first two stages of the Sea Launch Zenit-3SL, is designed for delivering payloads to inclined low Earth circular and elliptical orbits. This edition of the Land Launch User s Guide addresses a representative set of missions and launch services that can be cost-effectively accomplished with these two companion launch systems. Potential users are invited to complete the User Questionnaire (Appendix A) and return it to the address indicated therein. Rev. Initial Release HPD

24 Land Launch User s Guide Section 1 Figure 1-1. The Proven Sea Launch System Provides a Solid Foundation for Land Launch 1-2 HPD Rev. Initial Release

25 Land Launch User s Guide Section 1 What the system includes The standard Land Launch system comprises: The three-stage Zenit-3SLB launch vehicle The two-stage Zenit-2SLB launch vehicle The Zenit launch complex at Baikonur cosmodrome, and approved downrange stage and fairing impact zones Support facilities at Baikonur cosmodrome for ground processing of the spacecraft and launch vehicle including fueling, check-out, ILV assembly and launch Transportation equipment including rolling stock and thermostating systems for moving people and hardware between locations at Baikonur cosmodrome Tracking, meteorological and communications assets A thorough summary of Land Launch interfaces, operations, services and facilities is provided herein as Appendix B: Land Launch Standard Offerings and Options. Advantages to the Customer Land Launch provides: The most mature flight hardware in its payload class, closely derived from the proven Sea Launch configuration. Maturity is greatest with the Land Launch upper stage, the venerable Block DM, which is the most experienced and reliable upper stage in any payload class with continuous service since 1974 on more than 220 missions with a demonstrated reliability exceeding 97% The versatility of the Block DM, which has the capability for multiple restarts, long duration missions, roll and coast maneuvers, accurate orbital insertion and tightly controlled SC separation parameters of motion and attitude Dedicated launch, eliminating schedule and technical risk associated with co-passengers Existing, operational, proven ground facilities The teamwork and proven expertise of the Land Launch partners, which include the same core companies that work together on Sea Launch The responsiveness of the world s only dedicated commercial launch services family: Sea Launch and Land Launch Rev. Initial Release HPD

26 Land Launch User s Guide Section 1 Timeline It is expected that the first experience of a new spacecraft type can be integrated by eighteen months after contract signature. Repeat missions of the same spacecraft type can be integrated by twelve months after contract signature. Figure 1-2 provides a toplevel summary of the relative mission phases, while more extensive details on the mission flow process are presented in Section 8. Land Launch has the capability for conducting successive launches on 30-day centers at Baikonur. The vehicles are assembled off the launch complex, and the nominal time between vehicle roll-out and launch is on the order of twenty-eight hours, provided that spacecraft check-out on the pad does not exceed eighteen hours. Land Launch is dedicated to reducing the time required for integrating and launching spacecraft. Deviations from the standard flows may be accommodated on a case-by-case basis, particularly for two-stage missions. 18 Month Integration Timeline 18 mo 12 mo 6 mo 0 mo Sign Contract Mission Analysis Operations Planning New Mission S/C Delivery and processing Launch Ops 12 Month Integration Timeline 12 mo 6 mo 0 mo Sign Contract Mission Analysis Operations Planning Repeat Mission S/C Delivery and processing Launch Ops Figure 1 2. Land Launch Mission Integration Timelines Take Advantage of In-Place Sea Launch Procedures and Processes 1-4 HPD Rev. Initial Release

27 Land Launch User s Guide Section 1 Baikonur Cosmodrome The Zenit Launch Complex is located at the Baikonur Cosmodrome in Kazakhstan at 63 E, 46 N (Figure 1-3). Baikonur was a primary launch site for the Soviet Union and, in the post-soviet era, it continues to be the principal launch site for both the Russian and Ukrainian space industries. Many of the greatest events of the Space Age have occurred at Baikonur including launch of the world s first satellite in 1957, the first mission (unmanned) to the moon in 1959 and the first manned launch in Thousands of other launches have taken place at Baikonur over the ensuing decades up to the present day. In recent years Baikonur has become an important commercial spaceport. Since 1995 there have been more than fifty launches of satellites made outside the Commonwealth of Independent States (CIS). The cosmodrome is linked to major cities in the CIS by air, road and railway transport. The local area of the cosmodrome also has a developed road and railway network. The closest residential area is the city of Baikonur, located just south of the cosmodrome approximately 60 km from the Zenit Complex. Baikonur city lies on the north bank of the Syrdarya river. The Tyuratam settlement and the railway station of the same name (Kazakh railway) adjoin Baikonur. The Baikonur airport (Krainy) is linked to Moscow by regular and charter passenger flights and can accommodate most cargo and passenger airplanes. Spacecraft and cargo typically arrive via Yubileiny airport located within the cosmodrome itself. The Zenit space rocket complex is the newest operational launch facility at Baikonur, and conducted its first launch in It is located far from large populated areas, ensuring safety of launches and allowing for easy allocation of impact zones. Furthermore, its position on the eastern side of the cosmodrome enables greater flexibility with respect to launch azimuths than other local launch complexes can offer. The specific Baikonur cosmodrome facilities utilized by Land Launch are described more extensively in Section 6. Rev. Initial Release HPD

28 Land Launch User s Guide Section 1 Figure 1-3. The Baikonur Cosmodrome is Readily Accessible from Moscow 1-6 HPD Rev. Initial Release

29 Land Launch User s Guide Section Land Launch Organization Overview Land Launch contracts will be managed by the existing Sea Launch organization in Long Beach, California. Such co-location and shared use of resources and personnel is key to enabling Land Launch to provide a western interface to its customers that is comparable to that experienced with Sea Launch. Launch services out of Baikonur are obtained via subcontract from Sea Launch to Space International Services, Ltd (SIS). SIS is a limited liability company based in Moscow consisting of key Land Launch members from Ukraine and Russia, all of which also participate in Sea Launch missions. The Land Launch organizational structure is presented in Figure Program Management - Contracts and Legal Support - Insurance Interface - US Licensing Sea Launch Company, LLC Federal Space Agency (FSA) - CIS Licensing - CIS Governmental Interfaces Boeing - Sales & Marketing - Mission management - Quality and Technical Oversight - Hardware Acceptance Review - Customer Safety and Logistics Support Space International Services, Ltd. (SIS) Yuzhnoye SDO Yuzhmash PO KBTM TsENKI RSC Energia NPO Lavochkin - Launch Services - Mission Integration - Baikonur Operations Figure 1-4. The Land Launch Program Brings Together an Experienced Team Rev. Initial Release HPD

30 Land Launch User s Guide Section 1 Sea Launch Company, LLC Sea Launch Company, LLC, headquartered in Long Beach, California, is the world leader in commercial heavy-lift launch services with its highly successful and innovative ocean-based launch system. Sea Launch is a partnership comprised of Boeing, RSC Energia, SDO Yuzhnoye, the Kvaerner Group and PO Yuzhmash. Additional information related specifically to Sea Launch can be found in the Sea Launch User s Guide and on the corporate website at: For the Land Launch service, Sea Launch responsibilities include: Management of the overall endeavor Contracts & legal management Insurance and financing interfaces US licensing and US governmental interfaces The Boeing Company The Boeing Company is enlisted by Sea Launch to provide the following capabilities on Land Launch missions: Mission management Payload integration support Hardware quality reviews Overall technical and quality oversight Satellite safety assessments and customer logistics support Marketing & sales Space International Services, Ltd Space International Services, Ltd (SIS) is a company comprised of SDO Yuzhnoye, PO Yuzhmash, RSC Energia, the Center for Ground Space Infrastructure Operations (TsENKI) and the Design Bureau of Transport Machinery (KBTM). Its office is in Moscow, Russia. SIS direct responsibilities include: Launch services and Baikonur operations Supplier management (all CIS and launch hardware suppliers) CIS licensing and regulation Third party insurance 1-8 HPD Rev. Initial Release

31 Land Launch User s Guide Section 1 SDO Yuzhnoye SDO Yuzhnoye is the leading Ukrainian aerospace organization with vast experience in the design and development of launch vehicle technology. The company is established under the laws of Ukraine, with its principal place of business in Dnepropetrovsk, Ukraine. The SDO Yuzhnoye team along with that of PO Yuzhmash has conducted hundreds of successful launches from Baikonur (Fig. 1-5). Additional information related to SDO Yuzhnoye can be found on the website at: SDO Yuzhnoye performs the following work in support of the Land Launch program: Design and configuration management of the Zenit stages for both the Zenit-2SLB and Zenit-3SLB, as well as design support during their manufacturing Design of the Zenit-2SLB fairing and integrated launch vehicle (ILV) as a whole Systems engineering and integration Payload integration and mission analysis Technical management and participation in ILV processing and launch operations Rev. Initial Release HPD

32 Land Launch User s Guide Section 1 Figure 1-5. The Zenit Vehicle Reflects Five Decades of SDO Yuzhnoye and PO Yuzhmash Experience in Designing, Building and Operating Launch Vehicles PO Yuzhmash PO Yuzhmash is another leading Ukrainian aerospace enterprise having vast experience in the development and production of major launch vehicles. The company is incorporated under the laws of Ukraine and, like SDO Yuzhnoye, its principal place of business is in Dnepropetrovsk, Ukraine. PO Yuzhmash performs the following work in support of the Land Launch program: Manufacturing of the first two stages, and the Zenit-2SLB fairing ILV integration Participation in ILV processing and launch operations 1-10 HPD Rev. Initial Release

33 Land Launch User s Guide Section 1 Design Bureau of Transport Machinery (KBTM) The principal offices of KBTM are located in Moscow, Russia. In its fifty year history KBTM has designed numerous launch complexes including the Zenit complex at Baikonur (Figure 1-6), and also performs vehicle integration and launch operations. More than 900 orbital launches have been conducted from launch complexes built by KBTM. KBTM is a major subcontractor on the Sea Launch program responsible for ground support equipment maintenance including the transporter/erector. On Land Launch, KBTM will have overall responsibility for: ILV assembly area and launch complex ILV processing for launch Ground support equipment Launch operations Figure 1-6. KBTM Developed the ILV Transporter and Erector Equipment for both the Sea Launch and Land Launch Programs Rev. Initial Release HPD

34 Land Launch User s Guide Section 1 Center for Ground Space Infrastructure Operations (TsENKI) The principal offices of TsENKI are located in Moscow, Russia in close association with the Russian Space Agency. TsENKI is responsible for operation of the ground aerospace infrastructure facilities at Baikonur cosmodrome and downrange sites. TsENKI will have fundamental responsibility on Land Launch for: Russian launch licenses and third party insurance Recording, acquisition and processing of telemetry data Security and guard services Telecommunication services and communication systems Logistics support (propellant components and compressed gases) Securing cosmodrome support services during ILV processing and launch Processing operation and maintenance of impact zones Coordinating electromagnetic compatibility for ILV processing and launch operations 1-12 HPD Rev. Initial Release

35 Land Launch User s Guide Section 1 RSC Energia RSC Energia is the premier Russian space company. RSC Energia, developer of launch vehicles and propulsion systems, spacecraft, space stations, as well as manned and cargo modules, brings its legendary experience in space exploration and launch system integration to Land Launch. Energia is a joint stock company established under the laws of the Russian Federation, with its principal place of business in Korolev (near Moscow), Russia. Additional information related to RSC Energia can be obtained at: Energia has the responsibility on Land Launch for: Design and manufacture of the Block DM-SLB upper stage (Figure 1-7) Integration of the Ascent Unit comprising the Block DM-SLB, fairing, adapter and spacecraft Mission analysis support Launch operations support Customer support Figure 1-7. Energia s Block DM is the Most Successful and Most Proven Upper Stage Available Today Rev. Initial Release HPD

36 Land Launch User s Guide Section 1 NPO Lavochkin NPO Lavochkin is located in Khimki, Russia and has a distinguished history of achievement in the design, development and manufacture of aircraft, launch vehicle upper stages and spacecraft including many deep space missions to the moon, Venus and Mars. Lavochkin has also provided more than 100 fairings for various launch vehicles and will be providing Land Launch with a flight-proven fairing for the Zenit-3SLB. Federal Space Agency Land Launch enjoys the support of the Federal Space Agency which will be providing Land Launch with the use of its facilities at Baikonur, launch licensing and associated regulatory support including relations with other CIS governments on whose territory Land Launch activities will be conducted HPD Rev. Initial Release

37 Land Launch User s Guide Section 2 2. VEHICLE DESCRIPTION Overview Land Launch uses either two or three in-line, liquid oxygen (LOX) and kerosene stages. The three-stage Zenit-3SLB configuration (Figure 2-1) is used for medium-lift missions to medium and high, circular or elliptical orbits including GTO and GEO, as well as escape trajectories. The two-stage Zenit-2SLB configuration (Figure 2-2) is used for missions to low earth circular and elliptical orbits. Each configuration uses a different fairing. All elements of either configuration have extensive flight heritage. The principal components of the Land Launch vehicles are: Zenit Stage 1 Zenit Stage 2 Block DM-SLB upper stage (Zenit-3SLB configuration) Fairing and Payload Support Structure m (192.4 ft) Fairing 10.4 m (34.1 ft) Block DM-SLB Zenit Stage m (34.1 ft) Zenit Stage m (107.9 ft) Ø3.7 m (12.1 ft) Ø4.1 m (13.5 ft) Ø3.9 m (12.8 ft) Figure 2-1. The Zenit-3SLB Fairing 13.7 m (44.8 ft) Intersection Bay 0.35 m (1.1 ft) Zenit Stage m (34.1 ft) 57.4 m (188.3 ft) Zenit Stage m (107.9 ft) Ø3.9 m (12.8 ft) Figure 2-2. The Zenit-2SLB Rev. Initial Release HPD

38 Land Launch User s Guide Section 2 Design The Zenit first and second stages used on Land Launch are interchangeable with the Sea Launch first and second stages. They are manufactured by PO Yuzhmash in Ukraine, with design oversight provided by SDO Yuzhnoye. The Block DM-SLB third stage, used only on the Zenit- 3SLB, is closely adapted from the Block DM-SL used by the Sea Launch program (the differences are described in section 2.2) and is manufactured by RSC Energia in Russia. The fairing for the Zenit-3SLB is 4.1 meters in diameter and is manufactured by NPO Lavochkin in Russia. It was designed specifically for the Block DM and has an unblemished flight history dating to The Zenit-2SLB fairing is 3.9 meters in diameter and is manufactured by PO Yuzhmash. It was designed specifically for the two-stage Zenit configuration and has a flight history dating to The payload support structure for the Zenit-3SLB is provided by RSC Energia. It consists of a spacecraft adapter (SCA) typically procured from Saab Ericsson Space (937, 1194 or 1666 interfaces) and a transfer compartment manufactured by Energia. The payload support structure for the Zenit-2SLB is provided by SDO Yuzhnoye and will typically consist of a Saab SCA mounted on a truss manufactured by PO Yuzhmash. Unique interfaces and multi-satellite dispensers can also be provided if required. Zenit Flight History Block DM Flight History The original Zenit-2 was first launched in 1985 from Baikonur Cosmodrome. As of March 2004, it has completed 30 successful missions in 35 launch attempts. The Zenit first-stage booster also served as the strapon for the Energia launch vehicle (four per launch) and logged an additional eight successes in two flights in this capacity. The modified and improved Zenit-2S, the version used on Sea Launch, has flown twelve times as of February Land Launch also uses the Zenit-2S. From its introduction in 1974 through March 2004 the Block DM has completed 222 successful missions in 228 attempts in various versions, including eleven successes in eleven attempts for the Block DM-SL version used on Sea Launch, making it far and away the most proven, reliable and mature upper stage in the launch industry. Past missions have included GTO and direct insertion GEO for commercial and for government customers, high elliptical orbits, low and high circular orbits, dedicated launches and launches of multiple satellites, and escape trajectories (to Halley s Comet, Venus and Mars). 2-2 HPD Rev. Initial Release

39 Land Launch User s Guide Section 2 Flight Success Ratios Tables 2-1 and 2-2 list flight records for each of the three Zenit-3SL stages as of March 2004, as well as engineering reliability estimates. The closely-related Zenit-3SLB stages are expected to achieve identical reliability levels. The engineering reliability estimates account for: Extensive testing performed when modifications are made to flight hardware or ground support equipment Expected reliability growth, using statistics of other boosters using similar processes and procedures that were also built and launched in the former USSR An exhaustive failure analysis team that investigates any flight anomalies and implements measures to ensure that the anomalies never recur The Sea Launch mission assurance and audit process currently in place and operating at the factory level in Ukraine and Russia Table 2-1. Sea Launch Stages, Cumulative Flight History, All Related Configurations Stage Year Introduced Versions Flown Cumulative Flight Record Zenit Stage of 55 Zenit Stage of 44 Block DM of 228 Table 2-2. Reliability and Flight History, Sea Launch Configuration Only Stage Year Introduced Flight Record Reliability Estimate Zenit Stage of 12 Zenit Stage 2* of 11 } 98.0% Block DM-SL of % * The one Sea Launch failure (mission 3) occurred during second stage operation, but was not caused by the second stage and no design changes to the second stage resulted from the failure investigation. The failure cause was a fault in ground software that left open a second stage valve. Rev. Initial Release HPD

40 Land Launch User s Guide Section The Land Launch Zenit Design Heritage Land Launch uses the Sea Launch configuration of the Zenit, retaining the improvements and modifications that were made for Sea Launch to the heritage Zenit-2. SDO Yuzhnoye designed the original two-stage Zenit-2 during the late 1970s and early 1980s in response to requirements from the Soviet Ministry of Defense for a launch system that would be able to quickly and efficiently reconstitute military satellite constellations. Consequently, the design emphasizes robustness, ease of operation and fast reaction times, which are achieved through extensive automation. It incorporates state-of-the-art launch and processing technologies, developed by Land Launch partner KBTM, in contrast to systems developed during previous decades. A second intended use for the original Zenit-2 was manned launches to space station MIR (figure 2-3). Though ultimately it was never used for this purpose due to the break-up of the Soviet Union, in order to be man-rated, the Zenit was designed with a significant degree of internal redundancy and other features to ensure high reliability. Figure 2-3. Cosmonaut Access Tower at the Zenit Launch Complex 2-4 HPD Rev. Initial Release

41 Land Launch User s Guide Section 2 Changes Made for Sea Launch Significant configuration differences between the heritage Zenit-2 and the Sea Launch Zenit-2S, which are also retained on the Land Launch Zenit- 2SLB, are: New navigation system Next generation flight computer Increased performance due to mass reductions and an increase in second stage main engine thrust from 87 tonnes force to 93 tonnes force Avionics Just as on Sea Launch, the Land Launch Zenit contains its own complete complement of avionics for telemetry, guidance and navigation functions even when lifting an upper stage in a three-stage configuration. The onboard Sirius telemetry packages transmit telemetry data on separate RF links to existing ground stations located in Russia and, for sunsynchronous missions, to a remote ground station located on the Arabian Peninsula. For three-stage missions, these Zenit links are complemented by an independent set of data that is provided simultaneously by the Block DM-SLB telemetry system. Overall Specifications and Configurations Zenit specifications and performance parameters are shown in Table 2-3. Stage 1 and Stage 2 configurations are pictured in Figure 2-4. With propellant mass fractions exceeding 90%, the designs of both stages rank among the most structurally efficient in the world. In the case of the first stage, this is due in large part to the highly efficient RD-171M engine and the lack of strap-on boosters. The absence of strap-on boosters greatly simplifies pre-launch processing and is a major feature distinguishing Zenit from most other large launch systems. Without strap-ons, the stage structure is more efficient, ordnance count is reduced and overall reliability is enhanced by eliminating exposure to the failure of booster separation mechanisms or of the boosters themselves. Furthermore, the streamlined configuration lends itself to robust control margins during all phases of flight which enable the Zenit to fly through a broad range of wind and weather conditions, further ensuring on-time and on-target launch performance. Rev. Initial Release HPD

42 Land Launch User s Guide Section 2 Zenit Table 2-3. Land Launch Zenit Specifications Stage 1 Stage 2 Zenit-2SLB and -3SLB Zenit-2SLB Zenit-3SLB Burn Time s 300-1,100 s s Inert Mass 27,564 kg (60,768 lb) 8,367 kg (18,446 lb) 8,307 kg (18,314 lb) Fueled Mass 354,350 kg (781,200 lb) 90,854 kg (200,297 lb) 90,794 kg (200,164 lb) Fuel (kerosene) 90,219 kg (198,897 lb) 23,056 kg (50,829lb) Oxidizer (LOX) 236,567 kg (521,536 lb) 59,431 kg (131,022 lb) Length 32.9 m (108 ft) 10.4 m (34 ft) Diameter 3.9 m (12.8 ft) 3.9 m (12.8 ft) Engines Thrust (sea level) Thrust (vacuum) Specific impulse (sea level) Specific impulse (vacuum) One RD-171 (four thrust chambers) 740,000 kgf (1.63 million lbf) 806,400 kgf (1.78 million lbf) One RD-120 Main Engine One RD-8 Vernier Engine (four thrust chambers) Not applicable Main Engine: 93,000 kgf (205,028 lbf) Vernier Engine: 8,100 kgf (17,857 lbf) s Not applicable s Main Engine 350 s Vernier Engine s Attitude Control Nozzle gimbal deg Vernier engine nozzle gimbal + 33 degrees 2-6 HPD Rev. Initial Release

43 Land Launch User s Guide Section 2 Avionics bay Liquid oxygen tank Stage m (34 ft) Kerosene tank Main engine Separation plane SPST (4) Steering engine Interstage frame 3.9 m (12.8 ft) Stage m (108 ft) Liquid oxygen tank Solid-propellant separation thrusters (SPST) (4) Main engine nozzles (4) Kerosene tank Single turbopump View A-А А А Figure 2-4. Land Launch Uses the Same Zenit Stages that are Used on the Sea Launch Zenit-3SL Rev. Initial Release HPD

44 Land Launch User s Guide Section Zenit Stage 1 Overall Configuration The Land Launch Zenit Stage 1 (Figure 2-5) features an aluminum primary structure with integrally machined stiffeners, and environmentally-friendly LOX/kerosene propellants. The upper LOX tank fits in a concave depression at the top of the kerosene tank, and the LOX feed line runs through the middle of the lower tank. With a Zenit-2SLB gross lift-off mass of 450, ,000 kg, and a Zenit-3SLB gross lift-off mass of 462, ,000 kg, the 740,000 kgf produced by the first stage yields a very healthy ~1.6 takeoff thrust-to-weight ratio for both vehicles. Separation is achieved with four solid retro-rockets located at the base of the stage. The Land Launch Zenit first stage design is intentionally kept common to that of the Sea Launch stage. Both are manufactured on the same production line at PO Yuzhmash. Figure 2-5. PO Yuzhmash Achieves Significant Economies of Scale by Manufacturing Both Land Launch and Sea Launch Zenit Stages on the Same Production Line 2-8 HPD Rev. Initial Release

45 Land Launch User s Guide Section 2 RD-171M Engine The RD-171M engine (Figure 2-6), which powers Zenit Stage 1, burns liquid oxygen (LOX) and kerosene. It provides an impressive 740,000 kgf (1.6 million lbf) of thrust at sea level and is one of the most powerful rocket engines in the world, featuring advanced rocket engine technologies developed by leading Russian propulsion organizations. It was developed specifically for the Zenit, in parallel with the closely related RD- 170 that served as the strap-on booster for the Energia/Buran. An exhaustive test program consuming more than 200 test engines preceded first flight in the mid 1980 s. The four thrust chambers are fed by a single, vertically mounted turbopump, which in turn is powered by two gas generators feeding hot oxidizer-rich gas to a single turbine. Flight control is achieved by gimbaling the independently suspended combustion chambers, while the ability to throttle down to ~ 74 % of nominal full-engine thrust provides great flexibility in trajectory design J3-038 Figure 2-6. The RD-171M is the Most Powerful Liquid Rocket Engine Presently in Operation Rev. Initial Release HPD

46 Land Launch User s Guide Section Zenit Stage 2 Overall Configuration Like the first stage, the Zenit second stage (Figure 2-7) features an integrally stiffened aluminum construction and environmentally-friendly LOX/kerosene propellants. Propulsion is provided by an RD-120 main engine with steering provided by an RD-8 vernier engine fed from the same propellant tanks. The lower kerosene tank is toroidally shaped and surrounds the main engine, while the upper LOX tank is a domed cylinder. The stage is topped by an instrument compartment containing the avionics. The Sea Launch and Land Launch Zenit second stages, like the first stages, are manufactured on a common Yuzhmash production line, thereby benefiting from common inventory and Boeing quality oversight processes. The second stage generates 101,000 kg (222,887 lbs) of thrust (RD-120 and RD-8 engines combined). As on the first stage, separation is achieved with four aft-mounted solid retrorockets. Figure 2-7. The Second Stage s Toroidally-Shaped Fuel Tank Results in a Shorter, More Efficient Vehicle Structure J3-037R HPD Rev. Initial Release

47 Land Launch User s Guide Section 2 RD-120 Main Engine The 2 nd stage main engine is a single-chamber, fixed nozzle liquidpropellant rocket engine that uses LOX and kerosene to generate 93,000 kg f (205,028 lbf) of thrust. The RD-120 is throttled down to 78% of nominal full-engine thrust at the end of flight. The RD-120 was developed specifically for the Zenit launch system. RD-8 Vernier Engine The RD-8 vernier engine mounted in the aft end of Stage 2 provides three-axis attitude control. The RD-8 uses the same propellants and propellant storage system as the RD-120, with one turbo-pump feeding four gimbaling thrusters spaced around the outside of the RD-120. The RD-8 produces 8,100 kgf (17,900 lbf) of thrust, and was specifically developed for Zenit. The ability to modulate its operation from 65 to 900 seconds following main engine cut-off provides flexibility in mission design for Zenit-2SLB launches to a wide range of circular LEO orbits. Rev. Initial Release HPD

48 Land Launch User s Guide Section The Block DM-SLB Upper Stage Overall Configuration The Block DM-SLB (Figure 2-8) used on the Zenit-3SLB is closely derived from the Block DM-SL used on Sea Launch. It is a LOX/kerosene upper stage capable of igniting up to three times during a mission. Basic specifications are provided in Table 2-4. The basic structure of the Block DM-SLB is provided by the upper adapter together with an internal truss. The middle and lower adapters that enclose the stage are jettisoned before first ignition of the Block DM- SLB. Kerosene is contained in a toroidal tank connected by a truss to the upper adapter which encircles the turbopump of the 11D58M main engine. The spherical LOX tank and the avionics/payload truss are located above the kerosene tank, and also connect to the upper adapter. Two attitude control/ullage engines, which provide stabilization during coast periods, are located on the bottom of the kerosene tank. 11D58M Main Engine Attitude Control/Ullage Engines Avionics The Block DM-SLB upper stage is powered by the 11D58M engine, which operates on liquid oxygen and kerosene. Its carbon-carbon nozzle is gimbaled to provide pitch and yaw control during powered flight, with turbopump bleed gas used for roll control. Three axis stabilization and attitude control during coast periods, including continuous rolls, are provided by two attitude control/ullage engines using hypergolic propellants that are located on the aft end of the main engine kerosene tank, on either side of the main engine nozzle. The Land Launch Block DM-SLB uses the same avionics as the Sea Launch Block DM-SL, with the exception of differences in the telemetry system more suitable for launches from Baikonur using associated fixed and mobile Russian ground receiving stations. Changes Made for Sea Launch Significant configuration differences between the heritage Block DM and the Sea Launch Block DM-SL, which are also retained on the Land Launch Block DM-SLB, are: New navigation system Next generation flight computer The autonomous control system provided by the R&D and Production Center for Automation and Instruments Manufacturing (NPTs AP) the premier Russian avionics and space software company An extended nozzle and various mass reductions for performance improvement 2-12 HPD Rev. Initial Release

49 Land Launch User s Guide Section 2 Block DM-SLB Versus the Block DM-SL The main differences between the Block DM-SLB and the Block DM-SL are: The Block DM-SLB forward structural interfaces are made to be compatible with the Russian-made fairing and payload structure that are used on Land Launch, while the Block DM-SL forward interfaces are compatible with the Boeing-made payload unit hardware that is used on Sea Launch The single, large (and heavy) toroidal avionics bay on the Block DM-SL is replaced on the Block DM-SLB with several discrete avionics containers for a net reduction in launch mass Some sensors and harnesses are removed that are a legacy of early qualification flights and are no longer needed A deployable antenna and telemetry system are replaced with a lighter system also used on Zenit that features fixed antennas with two independent radio links An uplink command system and its antenna are removed One set of fuel tanks for the attitude control/ullage engines are removed. Previously, these tanks were routinely under-filled by the equivalent of one set of tanks. The LOX tank is pressurized with helium instead of an oxygen/helium mixture The minimum useable propellant criterion for the final re-start is lowered from 4000-kg to 1500-kg, by adding two 10-kgf thrusters to ensure settling prior to ignition An external heat radiator is removed with this function being assumed by the upper adapter structure Length 1 Diameter (primary) Table 2-4. Block DM-SLB Specifications Maximum Launch Mass 2, 3 (fueled) Maximum 3 Useable Propellant Reserve Thrust (vacuum) Note 1: Note 2: Note 3: 5.93 m (19.4 ft) 3.7 m (12.1 ft) 17,800 kg (39,240 lb) 14,580 kg (32,140 lb) 8,103 kgf (17,864 lbf) The fairing overlays 1.03 m (3.4 ft) of the length of the Block DM-SLB, as shown in Figure 2-8 Includes the lower and middle adapters, which are jettisoned prior to first burn of the Block DM-SLB Fuel is off-loaded for heavier payloads launching east (includes GTO missions), due to drop zone constraints Rev. Initial Release HPD

50 Land Launch User s Guide Section 2 Avionics Container Avionics/Payload Truss Coolant Piping Upper Adapter LOX Tank Middle Adapter Kerosene Tank Attitude Control/Ullage Engine Lower Adapter Main Engine Figure 2-8. Block DM-SLB (dimensions in millimeters) 2-14 HPD Rev. Initial Release

51 Land Launch User s Guide Section Zenit-3SLB Ascent Unit Components and Integration The Zenit-3SLB Ascent Unit (figure 2-9) consists of the spacecraft, Block DM-SLB, fairing and payload support structure (PSS). These elements are integrated in a Class 100,000 clean environment during ground processing. The PSS is comprised of an industry-standard spacecraft adapter typically procured from Saab Ericsson Space and a transfer compartment provided by RSC Energia. Payload Fairing (PLF) Spacecraft Spacecraft Interface Plane Payload Support Structure (PSS) PSS/BDM Interface PLF/BDM Interface Plane Stage 2/Stage 3 Block DM-SLB (BDM) Interface Plane Figure 2-9 Zenit-3SLB Ascent Unit (dimensions in millimeters) Payload Fairing Fairing Access Characteristics The payload fairing (PLF) provides environmental protection for the spacecraft from encapsulation in the payload processing facility through launch and ascent. The PLF for the Zenit-3SLB is based on the 17S72 fairing manufactured by NPO Lavochkin. It was designed specifically for the Block DM and has an unblemished flight record on Block DM missions dating to The fairing is a bi-conic, aluminum construction that is 10.4 meters (34.1 feet) in length by 4.1 meters (13.5 feet) in its primary diameter. Spacecraft interfaces provided by the PLF are described in further detail in Section 5. Once inside the PLF, physical access to the spacecraft is gained through fairing doors. Two doors are standard, one in each fairing half, up to 420 mm x 420 mm (16.5 inches x 16.5 inches) in size. Further information about access doors including allowable locations is provided in Section 5. Because there is no access tower at the Zenit launch pad, the customer/user can directly access their Land Launch payload(s) as late as 28 hours before launch, inside a clean enclosure at the Launch Vehicle Integration Facility. This capability improves opportunities for final adjustments, battery installation and other spacecraft-unique pre-launch operations. Rev. Initial Release HPD

52 Land Launch User s Guide Section 2 Conditioned Air Supply to the Fairing Fairing Thermal Protection Clean, conditioned air is provided to the payload fairing volume from encapsulation until launch including during transport between facilities. Flow rates, cleanliness, temperatures, humidity levels and other details of the clean air supply to the payload volume are provided in Section 4. The internal and external thermal insulation of the PLF nose cone protects the PLF structure against overheating and preserves acceptable thermal conditions for the spacecraft during ascent. Spacecraft environments are described in Section 4. Payload fairing jettison is constrained to ensure that the free molecular heating does not exceed the allowable limit defined in Section 4 and that the fairing elements land in pre-approved drop zones. Payload Support Structure The payload support structure for the Zenit-3SLB is provided by RSC Energia. It consists of a transfer compartment manufactured by Energia and an industry-standard spacecraft adapter (SCA) typically procured from Saab Ericsson Space Company (PAS937, PAS1194 or PAS1666) that interfaces with the spacecraft. Further details are provided in Section 5. Unique spacecraft base interfaces can normally be accommodated within the standard integration time span HPD Rev. Initial Release

53 Land Launch User s Guide Section The Zenit-2SLB Payload Unit Components and Integration The Zenit-2SLB payload unit (PLU), shown in Figure 2-10, consists of the spacecraft, fairing, intersection bay, interface truss and spacecraft adapter. These elements are integrated in a Class 100,000 clean environment during ground processing. Fairing Spacecraft Adapter Interface Truss Ø3.9 m (12.8 ft) Intersection Bay 13.6 m ( 44.8 ft) 14.0 m ( 45.9 ft) Figure 2-10 Zenit-2SLB Payload Unit Payload Fairing Fairing Access Characteristics The PLF for the Zenit-2SLB is based on the Zenit-2 fairing manufactured by PO Yuzhmash. It was designed specifically for the two-stage Zenit and has an extensive and unblemished flight record dating to The fairing is a mono-conic, aluminum construction that is meters (44.8 feet) in length by 3.9 meters (12.8 feet) in its primary diameter, and provides a 3.48 meter (11.4 feet) useable diameter. Spacecraft interfaces provided by the PLF are described in further detail in Section 5. Alternative and modified fairings are also available. Interested customers are encouraged to contact Boeing Launch Services for further information. Access doors up to 500 mm x 500 mm (19.7 inches x 19.7 inches) can be provided in the Zenit-2SLB fairing for this purpose. Additional information can be found in Chapter 5. Because there is no access tower at the Zenit launch pad, the customer/user can directly access their Land Launch payload(s) as late as 28 hours before launch, inside a clean enclosure at the Launch Vehicle Integration Facility. This capability improves opportunities for final adjustments, battery installation and other spacecraftunique pre-launch operations. Rev. Initial Release HPD

54 Land Launch User s Guide Section 2 Conditioned Air Supply to the Fairing Fairing Thermal Protection Clean, conditioned air is provided to the payload fairing volume from encapsulation until launch including during transport between facilities. Flow rates, cleanliness, temperatures, humidity levels and other details of the clean air supply to the payload volume are provided in Section 4. External thermal insulation protects the payload structure from overheating and the internal thermal insulation limits the interior payload fairing surface temperature. Payload fairing jettison is constrained to ensure that the free molecular heating does not exceed the allowable limit defined in Section 4 and that the fairing elements land in pre-approved drop zones. Intersection Bay The intersection bay serves to preserve the mating interfaces on the forward end of Zenit Stage 2 for the Block DM, thus maximizing inventory flexibility by allowing each stage 2 to be used on any Sea Launch or Land Launch configuration as needed. On Zenit-2SLB the intersection bay also provides a solid base for the payload support structure (truss and adapter) and enables full encapsulation of the spacecraft while in the payload processing facility, creating an enclosed payload volume for easy cleanliness and environmental control with a conditioned air supply. Spacecraft Adapters For dedicated launches of a single spacecraft on the Zenit-2SLB, Land Launch can provide the customer any of the available standard adapters manufactured by Saab Ericsson Space, or an adapter provided by SDO Yuzhnoye and PO Yuzhmash using their experience in developing, testing and manufacturing adapters and separation systems for past Zenit-2 missions (Fig. 2-11) and for other launchers produced by Yuzhnoye and Yuzhmash (Cyclone, Dnepr). Further information on interfaces is provided in Section 5. Unique Interfaces and Multi- Spacecraft Launches For spacecraft that have unique interface and separation requirements, Land Launch can examine other heritage spacecraft adapter designs, including those that incorporate bolt-type attachment and separation mechanisms. Yuzhnoye and Yuzhmash also have extensive experience designing and launching multi-spacecraft mechanisms on several launch systems. Unique spacecraft base interfaces, or multi-spacecraft dispensers, can normally be accommodated within the standard integration time span HPD Rev. Initial Release

55 Land Launch User s Guide Section 2 КА SC 2-я ступень РН LV Stage 2 O2062 Figure Zenit 2 Spacecraft Adapter Developed by Yuzhnoye and Yuzhmash Rev. Initial Release HPD

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57 Land Launch User s Guide Section 3 3. PERFORMANCE Overview The Land Launch vehicles, the Zenit-2SLB and the Zenit-3SLB, can deliver spacecraft to a broad set of orbits. These include low, medium and high Earth orbits (LEO, MEO and HEO), geosynchronous transfer orbits (GTO), highly elliptical orbits, direct geostationary insertion (GEO) and Earth escape trajectories. Data presented in this section is intended to enable prospective users to make preliminary performance assessments. Please contact Boeing Launch Services for a performance quote specific to your mission requirements. Characteristics of performance are covered in Sections 3.1 through 3.8, including: Launch Window Availability Launch Site and Accessible Orbits Generic Ascent Trajectories Mass Performance Coast Phase Maneuvers Injection Accuracy Spacecraft Separation Conditions Rev. Initial Release HPD

58 Land Launch User s Guide Section 3 Performance Ground Rules Performance data in this section is based on the following set of ground rules: Payload capability, defined in terms of Payload Systems Mass (PSM), consists of the combined mass of the separated spacecraft and the spacecraft adapter including wire harnesses. For preliminary planning of missions manifesting a single payload, spacecraft adapter (and harness) masses of 140 kg and 200 kg are assumed for the Zenit-3SLB and Zenit-2SLB respectively. The masses of dispensers for multiple payloads, typical for Zenit-2SLB missions to LEO, are application unique. The maximum PSM for Zenit-3SLB is 5,000 kg due to structural limitations. For Zenit-2SLB the structural limit is not a factor since it exceeds the vehicle s maximum performance. To achieve orbit within the desired accuracy, and perform Contamination and Collision Avoidance Maneuver (CCAM), sufficient propellant reserves are assured for each individual stage to account for all launch vehicle dispersions and possible ambient conditions at any time of day on any day of the year with at least 99.65% probability. The spacecraft is injected into orbit via trajectories that are consistent with existing, approved launch corridors and drop zones. At the time of fairing jettison, the free molecular heating (FMH) is less than 1,135 W/m 2, accounting for all launch vehicle dispersions and possible ambient conditions at any time of day on any day of the year. Orbital altitudes are specified with respect to an Earth radius of 6,378 km. The Zenit-3SLB uses its standard payload fairing that is 4.1 m in diameter and 10.4 m long. The Zenit-2SLB uses its standard payload fairing that is 3.9 m in diameter and m long. Mission-unique customer requirements that may affect performance (e.g. specific argument of perigee, restricted mission duration, ground station visibility, extended launch windows) are not factored. 3-2 HPD Rev. Initial Release

59 Land Launch User s Guide Section 3 Launch Window Availability The launch vehicle and associated ground systems can support a launch window any day of the year at any time of the day. Furthermore, inherent features of the Land Launch system enable it to provide the maximum flexibility to accommodate shifting satellite readiness dates with little or no perturbation to the launch schedules of other customers (Table 3-1). Minimal Turn-Around Time - the Zenit launch complex was designed for maximum throughput and minimum refurbishment between launches. The complex can support launches as little as 10 days apart. Factory output limits the theoretical launch rate to twelve per year, of which seven may be Zenit-3SLB. Robust Flight Hardware Both the Zenit and Block DM launch systems were designed to withstand environmental conditions at Baikonur. Heritage Hardware The Land Launch configurations are composed of heritage systems with as many as 220 flights to their credit. This maturity, combined with robust commit criteria, give Sea Launch and Land Launch the highest launch-on-time probability for heavy and medium lift launch services, respectively. All but two of twelve Sea Launch launches to date have taken place in the first second of the first launch window on the first attempt. Table 3-1. Launch Operational Features Dates Available year around Times Ambient Temperature Average Ground Winds (at 10m above ground surface) Pad Turn-around Time Between Launches Nominal Turn-around Time After Launch Scrub Maximum Annual Launch Rate (Factory Limited) Launch-on-Time Probability Available at any hour -29 C to +45 C (-20 F to +113 F) Zenit-2SLB: 20 m/s (45 miles/hour) Zenit-3SLB: 18 m/s (40 miles/hour) 10 Days 1 Day (if scrub precedes LV fueling) < 3 Days (scrub after LV is fueled) Twelve (of which no more than seven Zenit-3SLB) Zenit-2SLB: 98% Zenit-3SLB: 97% Rev. Initial Release HPD

60 Land Launch User s Guide Section Launch Site and Accessible Orbits Site Location The coordinates for the Zenit Launch Complex are: latitude = 46 o North, longitude = 63 o East. The currently approved launch azimuths available from this complex, as constrained by overflight and drop zone considerations, are shown below in Table 3-2 and Figure 3-1. Table 3-2. Zenit Launch Azimuths and Inclinations from Baikonur Azimuth Inclination of Initial Orbit 64.2º 51.4º 35.0º 63.9º 194.2º 98.8º For special cases, arrangements can be made to open a corridor and allocate drop-zones for the launch azimuths of A o = 82.1 o (i = 46.2 o ) and A 0 =178.8º (i=88.1º). Approval of new launch corridors for Land Launch is eased by its use of environmentally-friendly fuels. Approved Azimuth Potential Azimuth A=35.0 0, i= A=64.2 0, i= Baikonur A=82.1 0, i= A= , i= A= , i= Figure 3-1. Flight Corridors for Land Launch from Baikonur 3-4 HPD Rev. Initial Release

61 Land Launch User s Guide Section 3 Accessible Orbits Table 3-3 shows the orbit inclinations (i) that can be reached by Land Launch from its three approved launch corridors. LEO orbit inclinations several degrees different from the three approved launch corridors can be obtained by cross-range yawing maneuvers ( doglegs ) of the second stage commencing after fairing jettison. Such maneuvers are generally associated with missions provided by the Zenit-2SLB, where LEO is the final destination. For Zenit-3SLB missions involving higher orbits in which the desired inclination differs from the three approved corridors, it is typically most efficient for plane changes to be carried out primarily by the Block DM-SLB third stage. In these cases, the first two stages usually perform a direct ascent into a parking orbit inclination coinciding with one of the approved corridors. Orbit Type LEO MEO, HEO, GTO, Elliptical, escape trajectories Table 3-3. Accessible Orbits on Land Launch Accessible Usual Plane Change Vehicle Inclinations Method 46.2 < i < 71 o 84.0 < i < 105 o Zenit-2SLB Second Stage Yaw 0.0 < i < 110 o Zenit-3SLB Third Stage Perigee, Apogee or Post- Perigee Burn (mission-specific) Performance losses due to plane changes are highly sensitive to a variety of mission parameters. Consequently, prospective Land Launch customers are encouraged to contact Boeing Launch Services for a performance estimate that is specific to their needs. Rev. Initial Release HPD

62 Land Launch User s Guide Section Ascent Trajectory Generic Zenit-3SLB GTO Mission Mission Profile For GTO missions the Zenit-3SLB flies a classic three-burn Block DM mission profile (Figure 3-2) using the approved corridor and drop zones at A o =64.2, i=51.4 (Figure 3-3). Block DM-SLB second burn Injection into transfer orbit Block DM-SLB first burn Injection into parking orbit Transfer orbit HP = 200 km HA = 35,950 km Inc = 48.6 deg Zenit Stage I/II Parking orbit HP = 180 km HA = 417 km Inc = 51.4 deg Block DM-SLB third burn Injection into target transfer orbit Target transfer orbit HP = 4,100 km HA = 35,786 km Inc = 23.2 deg Figure 3-2. Land Launch Uses the Proven Three-Burn Block DM Mission Profile From Baikonur for GTO Launches (orbit parameters correspond to PSM=3600 kg) 3-6 HPD Rev. Initial Release

63 Land Launch User s Guide Section 3 Launch point Fairing L=1924 km First Stage L= 884 km Second Stage L=6850 km Figure 3-3. Approved Land Launch Ground Track and Drop Zones for GTO Missions Stage 1 Flight The Zenit first stage provides the thrust for the first 149 seconds of flight. The roll maneuver begins at 10 seconds after launch. During the final seconds of its burn the engine is throttled to limit the maximum axial acceleration. The approved drop zone for the separated first stage is at distance of approximately 884 km from the launch point, within the Republic of Kazakhstan as shown in Figure 3-3. Throughout this phase of the mission, telemetry is received by ground stations within Baikonur cosmodrome. Rev. Initial Release HPD

64 Land Launch User s Guide Section 3 Stage 2 Flight The Zenit Stage 2 vernier engine ignites just prior to first stage separation. Upon first/second stage separation, the first stage solid retrorockets fire and second stage main engine ignition occurs. The second stage main and vernier engines continue to operate in tandem for the next five minutes of flight. After second stage main engine cut-off, the vernier engine continues to function for 75 seconds to provide attitude control up through second/third stage separation. Payload fairing jettison occurs at approximately 320 seconds into flight (175 seconds into second stage operation) with the drop zone located in Siberia approximately 1924 km downrange of the launch site. At this point the free molecular heating rate has dropped to below 30 W/m 2, well below the industry norm of 1,135 W/m 2. Cross-range yaw maneuvers by the second stage, if required, take place after fairing separation. Telemetry coverage during second stage flight is typically provided by ground stations at Baikonur cosmodrome, and at Krasnoyarsk in Russia. The second stage drop zone is located within the neutral waters of the Pacific Ocean at a downrange distance of 6850 km. 3-8 HPD Rev. Initial Release

65 Land Launch User s Guide Section 3 Block DM-SLB Powered Flight At approximately 65 to 90 seconds after second stage main engine shutdown and an altitude of 180 to 400 km, the second stage vernier engine shuts down. This event is quickly followed by second/third stage separation and the subsequent jettison of the middle adapter surrounding the Block DM-SLB. The Block DM-SLB can perform one to three burns. For most multipleburn missions, including the generic three-burn GTO mission described here, the initial burn establishes a stable parking orbit, begins approximately ten seconds after separation of the second stage and lasts approximately 200 seconds, with telemetry coverage provided from Krasnoyarsk. The Block DM-SLB then begins a coast in the parking orbit lasting about 64 minutes. Attitude control during Block DM-SLB coast phases is provided by its two attitude control/ullage engines. The second Block DM-SLB burn occurs at the first ascending node of the parking orbit, over the Atlantic Ocean, to transfer to an intermediate elliptical orbit with a synchronous or super-synchronous apogee as dictated by customer requirements and the capabilities of the satellite platform. Ignition starts at approximately 75 minutes after launch and typically continues for approximately 6 minutes, with telemetry coverage provided by a mobile receiving station. After a 5-hour coast the Block DM-SLB and payload reach GTO apogee, where a third burn is performed to optimize the delivery orbit by raising perigee and reducing inclination. Telemetry coverage during the third burn is simplified by the altitude at which it occurs, and is typically provided by multiple sites located at Moscow, Baikonur, Krasnoyarsk and elsewhere. The target injection orbit for a payload mass of 3600 kg features a perigee of 4100 km, an apogee of km and inclination of 23.2, resulting in a velocity shortage of 1500 meters/second required to achieve GEO. Payloads lighter than 3600 kg are delivered to orbits requiring progressively less than 1500 meters/second delta-velocity to GEO to the point that payloads weighing 1,600 kg and less are inserted directly into GEO, a mission that the Block DM family has already performed more than one hundred times. Spacecraft separation conditions and post-separation events including collision avoidance maneuvers are described in Section 3.8. Rev. Initial Release HPD

66 Land Launch User s Guide Section 3 Flight Timeline Table 3-4 provides a typical sequence of events for a representative threeburn Zenit-3SLB mission to GTO for a 3600-kg payload. Event timing is only slightly dependent on payload mass. Apart from spacecraft separation, variation (dispersion) of any planned event timing for a nominal mission is typically within 15 seconds from the reference sequence. Table 3-4. Flight Timeline GTO Mission by the Zenit-3SLB with Three Burns of the Block DM-SLB Time [seconds] Event 0 Ignition 3.9 Liftoff 12 Begin pitch over 14 Roll to launch azimuth 59 Maximum dynamic pressure 115 Maximum axial acceleration 115 to 132 Stage 1 engine throttle down to 74% 144 Stage 2 vernier engine ignition 147 Stage 1 engine shutdown 149 Stage 1 separation 154 Stage 2 main engine ignition 319 Payload fairing jettison 432 Stage 2 main engine shutdown 507 Stage 2 vernier engine shutdown 508 Stage 2 separation 509 Block DM-SLB middle adaptor jettison 517 Block DM-SLB main engine ignition #1 707 Block DM-SLB main engine shutdown # Block DM-SLB main engine ignition # Block DM-SLB main engine shutdown # Block DM-SLB main engine ignition # Block DM-SLB main engine shutdown #3 Mission-Specific Spacecraft separation 3-10 HPD Rev. Initial Release

67 Land Launch User s Guide Section 3 Ground Track Figure 3-4 presents the predicted ground track of injection for a generic, representative Zenit-3SLB three-burn GTO mission. Launch point 1-st ignition of DM-SLB ME 2-nd ignition of DM-SLB ME LV operation phase 3-rd ignition of DM-SLB ME Figure 3-4. Injection ground track for a generic Zenit-3SLB GTO mission. Rev. Initial Release HPD

68 Land Launch User s Guide Section Ascent Trajectory Generic Zenit-2SLB Mission to 51.6 o LEO Stage 1 Flight The two-stage Zenit-2SLB is optimized for LEO missions inclined at 51.6 o, including potential flights to the International Space Station (ISS). For such missions, the Zenit Stage 1 uses the same approved launch corridor and drop zone that is used for GTO missions, along launch azimuth 64.2 o (inclination 51.4 o ). Liftoff occurs 3.9 seconds after ignition, upon release of the hold-downs. The roll maneuver begins at 10 seconds into flight. Main engine thrust is provided for the first seconds of flight, and the engine is throttled during its last seconds of operation in order to limit maximum axial acceleration. The drop zone for the first stage is 884 km down range from the launch point, within the Republic of Kazakhstan. Throughout this phase of the mission, telemetry is received by ground stations within Baikonur cosmodrome. Stage 2 Flight The Zenit Stage 2 steering engine ignites prior to first stage separation. Upon first/second stage separation, the first stage solid retrorockets fire and second stage main engine ignition occurs. The main engine and vernier engine continue to operate in tandem for the next four minutes of flight. Fairing jettison occurs at about 295 seconds of flight (150 seconds into second stage operation), consistent with the approved drop zone located in Siberia approximately 1924 km downrange of the launch site. At this point the free molecular heating rate has dropped to below 30 W/m 2, well below the industry norm of 1,135 W/m 2. After fairing jettison, the second stage performs a cross-range yaw maneuver to adjust the inclination to 51.6 o. After second stage main engine cut-off, the vernier engine continues to function for an additional 500 seconds (as long as 890 seconds on other missions) to provide attitude control up through payload separation. Throughout this phase of flight, telemetry is received by the ground stations within Baikonur and Krasnoyarsk. Spacecraft separation conditions and post-separation events including collision avoidance maneuvers are described in Section HPD Rev. Initial Release

69 Land Launch User s Guide Section 3 Flight Timeline Table 3-5 provides a typical sequence of events for a representative Zenit-2SLB mission that delivers 12,000 kg to a 51.6 o -inclined, 400 km low Earth orbit, i.e. one compatible with ISS access. Table 3-5. Flight Timeline Zenit-2SLB ISS Mission Time [seconds] Event 0 Ignition 3.9 Liftoff 10 Begin roll maneuver 11 Begin pitch over 14 Roll to launch azimuth 60 Maximum dynamic pressure 113 Maximum axial acceleration 113 tо 132 Stage 1 engine throttle to 50% 145 Stage 2 vernier engine ignition 147 Stage 1 engine shutdown 149 Stage 1 separation 155 Stage 2 main engine ignition 295 Payload fairing jettison 397 Stage 2 main engine shutdown Stage 2 vernier engine shutdown Spacecraft separation pyrotechnic firing Solid-propellant retro rocket burn Rev. Initial Release HPD

70 Land Launch User s Guide Section 3 Flight Profile Figure 3-5 graphically portrays the flight profile defined in Table 3-5, along with other key trajectory events and parameters. Fairing Jettison Time=295s Altitude=173 km FMH 30 W/m 2 Stage 2 MECO Time=397 s Altitude=400 km SC Separation Time=894 s Altitude=400 km Stage 1 Separation Time=149 s Altitude=74 km Max Acceleration Time=113 s Accel=4.06 g Maximum Q Time=60 s Q=5370 kgf/m 2 Stage 1 Impact Range=884 km Fairing Impact Range=1924 km Figure 3-5. Typical Ascent Profile to the International Space Station Orbit at 51.6 o with Payload Mass kg 3-14 HPD Rev. Initial Release

71 Land Launch User s Guide Section 3 Ground Track Figure 3-6 presents the predicted ground track for a Zenit-2SLB mission to the International Space Station Stages 1 and 2 Operation Figure 3-6. Flight Ground Track for a Zenit-2SLB Mission to 51.6 o LEO Rev. Initial Release HPD

72 Land Launch User s Guide Section Payload Capability Three Stage Zenit-3SLB Geosynchronous Transfer Orbit The Land Launch Zenit-3SLB is a medium-lift vehicle to GTO. Employing three burns of the Block DM-SLB, it can deliver payloads weighing 3.6 metric tons to a GTO featuring a high perigee and reduced inclination, requiring 1500 m/s in additional velocity to attain geostationary or geosynchronous orbit (GEO). Performance improves rapidly for lighter satellites because correspondingly less fuel is offloaded from the Block DM-SLB to meet a second stage drop zone constraint. Table 3-6 and Figure 3-7 show the GTO payload capability. Table 3-6. Zenit-3SLB Payload Capability to GTO Delta-V to GEO [meters/second] Inclination [degrees] Perigee Altitude [kilometers] Payload Systems Mass [kilograms] ,786 1,600 1, ,430 2,830 1, ,100 3,600 1, ,120 4,120 Notes and Assumptions: Apogee altitude of 35,786 km Three burns of the Block DM-SLB Mission duration approximately 6.6 hours Payload Mass, kg Delta V to target orbit, m/s Figure 3-7. Zenit-3SLB Payload Capability to GTO 3-16 HPD Rev. Initial Release

73 Land Launch User s Guide Section 3 MEO, HEO, Circular and Elliptical Orbits The Land Launch Zenit-3SLB is a heavy lift vehicle to Middle Earth and High Earth (MEO and HEO, respectively) circular and elliptical orbits that coincide with its approved launch corridors, as shown in Tables 3-7 and 3-8 and in Figures 3-8 and 3-9. MEO, HEO and elliptical orbits at other inclinations can also be obtained, typically with an additional burn of the Block DM-SLB, at a cost in performance that varies with altitude and the extent of plane change required. LEO (altitude<1000 km) and low-perigee elliptical orbits are more optimally performed by a Zenit-2SLB, as shown in a later section of this chapter. Customers are encouraged to contact Boeing Launch Services for a specific performance quotation. Table 3-7. Zenit-3SLB Performance to Circular Orbits Payload Capability [kg] Height [km] Inclination Inclination Inclination 51.4 o 63.9 o 98.8 o 1, , , , , Note: Two burns of the Block DM-SLB main engine Payload Mass, kg i=51,4 i=63,9 i=98, Orbital Height, km Figure 3-8. Zenit-3SLB Performance to Circular Orbits Rev. Initial Release HPD

74 Land Launch User s Guide Section 3 Table 3-8. Zenit-3SLB Performance to Elliptical Orbits Apogee Payload Capability [kg] Height [km] Inclination 51.4 o Inclination 63.9 o Inclination 98.8 o 10, , , , , , , Assumptions: Single Block DM-SLB burn Perigee altitude of ~200 km Payload Mass, kg Height of target orbit apogee, km i=51,4 i=63,9 i=98,8 Figure 3-9. Zenit-3SLB Performance to Elliptical Orbits (Perigee 200 km) 3-18 HPD Rev. Initial Release

75 Land Launch User s Guide Section 3 High-Energy and Earth-Escape Trajectories Table 3-9 and Figure 3-10 show the Zenit-3SLB payload capability to high-energy orbits and Earth escape. These are presented as a function of C3 (velocity-at-infinity squared). Table 3-9. Zenit-3SLB High-Energy and Earth Escape Payload Capability C 3 [km 2 /s 2 ] Payload Capability [kg] Notes and Assumptions: Inclination = 51.4 o Perigee altitude = km Single Block DM-SLB burn Payload Systems Mass, kg C 3, km 2 /s 2 Figure Zenit-3SLB High-Energy and Earth Escape Payload Capability Rev. Initial Release HPD

76 Land Launch User s Guide Section Payload Capability - Two Stage Zenit-2SLB Circular LEO Orbits Table 3-10 and Figure 3-11 present Zenit-2SLB payload performance as a function of both circular orbit altitude and inclination. Table 3-10 Zenit-2SLB Payload Capability for Circular Low Earth Orbits Payload Mass [kg] Altitude [km] Inclination 51.4º Inclination 63.9º Inclination 98.8º ,920 13,330 10, ,940 12,410 9, ,930 11,500 8, ,890 10,550 7, ,820 9,570 6, ,730 8,560 5, ,630 7,550 4, ,530 6,510 3,940 1,000 5,420 5,480 3,320 1,100 4,660 4,560 2,920 1,200 4,250 4,190 2,530 1,300 3,810 3,750 2,320 1,400 3,390 3,310 2,030 1,500 2,930 2,340 1, Payload Mass (kg) º 63.9º 98.8º Circular Orbit Height (km) Figure Zenit-2SLB Payload Capability for Circular Low Earth Orbits 3-20 HPD Rev. Initial Release

77 Land Launch User s Guide Section 3 Elliptical Orbits Table 3-11 and Figure 3-12 define the performance parameters for the two-stage Zenit-2SLB to various elliptical earth orbits. Table Zenit-2SLB performance to Elliptical Orbits Payload Mass [kg] Apogee [km] Inclination 51.4º Inclination 63.9º Inclination 98.8º , , , , , , Note: Perigee altitude = 200 km Payload mass (kg) º 63.9º 98.8º Apogee altitude (km) Figure Zenit-2SLB Performance to Elliptical Orbits (Perigee 200 km) Rev. Initial Release HPD

78 Land Launch User s Guide Section Coast Phase Attitude Maneuvers Zenit-3SLB During coast phases the Block DM-SLB control system can provide three axes pointing (pitch, yaw and roll) with accuracy up to ±3 deg in all three axes. The control system of the Block DM-SLB, unlike other versions of the Block DM, can also provide continuous roll around the longitudinal axis or one of the lateral axes at a rate up to 5 degrees per second. Forty minutes of any coast phase are nominally reserved for Block DM-SLB attitude maneuvers Zenit-2SLB Zenit-2SLB missions do not feature extended coasts. Stage 2 operation immediately succeeds stage 1 operation, and payload separation occurs between 0.3 and 5 seconds after cut-off of the second stage vernier (steering) engine HPD Rev. Initial Release

79 Land Launch User s Guide Section Injection Accuracy Tables 3-12 and 3-13 show 3σ orbital injection accuracy of the Land Launch family of vehicles to representative orbits. Table Land Launch Zenit-2SLB and Zenit-3SLB Provide Accurate Orbital Insertion Orbital Parameter Zenit-2SLB Zenit-3SLB Circular (1) Circular (2) Circular (3) GTO (4) Altitude [km] ± 8 ± 9 ± 25 - Perigee [km] ± 40 Apogee [km] ± 100 Inclination [deg] ± 0.04 ± 0.07 ± 0.06 ± 0.1 Longitude of Ascending Node [deg] ± 0.1 ± 0.07 ± 0.2 ± 0.3 Perigee Argument [deg] ± 0.2 Period [sec] ± 3.5 ± 4.5 ± 45 (1) 400 km x 400 km, inclination = 51.6 (2) 600 km x 600 km, inclination = 98 (3) 10,000 km x 10,000 km, inclination = 51.4 (4) 4,000 km x 35,786 km, inclination = 23 Table The Zenit-3SLB Also Provides Accurate Direct GEO Insertion Orbit Type Orbital Altitude Inclination Period Geostationary ± 200 km ± 0.2 deg. ± 450 s Rev. Initial Release HPD

80 Land Launch User s Guide Section Spacecraft Separation and Post-Separation Events Zenit-3SLB Separation Event Spacecraft separation typically occurs minutes after the final Block DM-SLB main engine shutdown. This allows for reorientation to the required spacecraft separation attitude. Separation Capabilities The separation system provides a relative velocity between the Block DM-SLB and the spacecraft, typically on the order of 0.3 meters/second. The separation springs can provide a straight push-off or a transverse angular rate. Attitude and attitude rate accuracy depend heavily on spacecraft mass properties and spin rate, and may be assumed to be degrees and degrees/second in all three axes for a non-spinning separation (2.3σ). The Block DM-SLB attitude control system can provide a longitudinal spin rate up to 5 degrees per second if desired. For spacecraft requiring a transverse spin at separation, this may be provided up to 2 degrees per second within +/- 0.5 degrees per second about each axis. CCAM After spacecraft separation, the Block DM-SLB performs a Collision and Contamination Avoidance Maneuver (CCAM), which prevents future contact with the spacecraft. The timing of this maneuver is determined for the specific mission. The Block DM-SLB then vents all residual propellant and gasses, and depletes any remaining charge in its batteries. State Vector Delivery The state vector at time of spacecraft separation may be delivered to the customer minutes after the event. The format of the state vector, means of its delivery and the parameters of the spacecraft injection orbit are agreed in advance between the parties. The time of delivery of data can be updated for the specific mission HPD Rev. Initial Release

81 Land Launch User s Guide Section Zenit-2SLB Separation Event Separation begins between 0.3 and 5 seconds after shutdown of the second stage vernier (steering) engine. Separation Capabilities The Zenit-2SLB employs a typical three-axis stabilized method for payload separation along the second stage s longitudinal axis. The actual separation is initiated by the firing of pyrotechnic ordnance charges in the spacecraft attachment assembly. The separation impulse to the spacecraft is typically provided by springs in the separation system. Nearly simultaneously, solid propellant retro-rockets on the aft end of the second stage are fired, adding to the relative separation velocity. Launch vehicle stabilization errors at the moment of spacecraft separation command generation can be kept within +/- 2 degrees for pitch and yaw and within +/- 1 degree for roll. Angular velocities at release can be kept within +/- 1.5 degrees/sec for all three axes. Table 3-14 presents typical parameters for payload motion after separation in the case of a single spacecraft, while Table 3-15 presents similar data for missions involving multiple payloads with individual masses that exceed 500 kg. Table 3-14 Typical Spacecraft Motion After Separation - Single Payload Parameter Value Relative separation velocity 2.8 m/s Spacecraft angular rate around any of its axes 2.5 deg/s Spacecraft attitude error ± 2 deg Table 3-15 Typical Spacecraft Motion After Separation - Multiple Payloads (each > 500 kg) Parameter Relative separation velocity Spacecraft angular rate around any of its axes Spacecraft attitude error Value 0.3 m/s 4.0 deg/s ± 5 deg Rev. Initial Release HPD

82 Land Launch User s Guide Section 3 CCAM for second stage Collision avoidance is achieved by firing four solid-propellant retrorockets on the aft end of the second stage for a burn time on the order of 0.5 to 1.1 seconds, slowing the second stage and moving it out of the spacecraft orbit. After a delay, the oxidizer tank is vented. State Vector Delivery The timing of state vector delivery depends on the mission profile as well as the location and the availability of ground stations. For a typical ascent to 51.4 o, it is possible to arrange for delivery of such data to the customer between 35 and 50 minutes after payload separation HPD Rev. Initial Release

83 Land Launch User s Guide Section 4 4. SPACECRAFT ENVIRONMENTS Overview This section describes the major environments to which the spacecraft is exposed from the time of its arrival at Baikonur cosmodrome until its separation from the launch vehicle during flight. These environments and conditions include: Structural loads Thermal Random vibration Humidity Acoustics Pressure venting Shock Contamination Electromagnetic radiation Unless otherwise noted, the payload environments presented in this section are common to both the three-stage Zenit-3SLB and the two-stage Zenit-2SLB. Ground and Flight Environments Those levels associated with ground handling and transportation address the period from the arrival of the spacecraft at Baikonur until Stage 1 ignition and liftoff. Those levels designated as flight cover the subsequent period from liftoff command through spacecraft separation. Reference Coordinate System The coordinate system used in this section is shown in Figure 4-1. During transfer of the spacecraft in its shipping container from the airport to the Payload Processing Facility (PPF), the +X axis coincides with the direction of travel. At other times, X coincides with the longitudinal axis of the launch vehicle. The Y axis is vertical during horizontal ground operations. Rev. Initial Release HPD

84 Land Launch User s Guide Section 4 Y Vertical Z Lateral X Longitudinal Figure 4-1. Reference Coordinate System Used for Defining Spacecraft Environments Environmental Monitoring Land Launch monitors and records spacecraft environments as specified in the Spacecraft-LV Interface Control Document and documents these results in the post-flight report to the customer. Typically, this includes: flight environments (accelerations, acoustics, shock, fairing thermal conditions during ascent, pressure decay, etc.); the temperature, humidity and cleanliness levels of the spacecraft processing and encapsulation areas while the spacecraft is present; the temperature, humidity and cleanliness levels of the conditioned air provided to the fairing with the spacecraft inside; accelerations experienced during all phases of ground processing, after the spacecraft has been removed from its shipping container The responsibility normally resides with the customer to monitor the spacecraft environment (including accelerations) until it is unloaded from its shipping container at the cosmodrome. 4-2 HPD Rev. Initial Release

85 Land Launch User s Guide Section Structural Loads Overview Design reference structural loading environments on spacecraft primary and secondary structures are defined here for: ground transportation and handling flight spacecraft sinusoidal vibration testing Spacecraft compliance requirements related to these environments are presented in Section 7. Quasi-Static Load Factors, Ground Handling and Transportation Design reference maximum acceleration levels during ground transportation, handling and processing are defined in Table 4-1. The quasi-static accelerations levels are shown for the spacecraft center of gravity while in a horizontal orientation. These accelerations can be applied simultaneously in the longitudinal, lateral and vertical directions (the axis X coincides with the velocity vector). During erection of the launch vehicle to a vertical position on the launch pad, the maximum acceleration of the spacecraft center of gravity is 1.5 g. In the course of combined operations the spacecraft briefly transitions through various vertical orientations, during which the respective axial accelerations are maintained within those limits already specified. The maximum rate of angular acceleration (about any axis) during crane lifts of the ILV is radians per seconds squared. Table 4-1. Maximum Quasi-Static Accelerations During Ground Operations Spacecraft Processing Operation Acceleration [g] Safety X Y Z Factor Transfer from the airport to the PPF ± ± 1.0 ± Horizontal Combined Operations (from mating with the launcher in the PPF through completion of launcher assembly in Area 42) Zenit-3SLB Zenit-2SLB ± 0.5 ± ± ± 0.2 ± 0.4 ± Launcher on-loading and off-loading (crane lifts in Area 42) ± ± 0.2 ± Roll out and erection on the launch pad ± ± 0.2 ± Rev. Initial Release HPD

86 Land Launch User s Guide Section 4 Quasi-Static Load Factors, Flight From liftoff through spacecraft separation, the spacecraft is subjected to quasi-static steady-state and low-frequency dynamic accelerations. Figure 4-2 provides the design reference accelerations for critical loading events. These accelerations are applied at the spacecraft center of gravity and are intended for preliminary design only. Determining the ability of specific spacecraft primary and secondary structures to withstand the dynamic loading events during flight requires a coupled loads analysis (CLA), which will be performed for each mission. When generated and verified, CLA results supersede the generic quasi-static accelerations provided in Figure 4-2. (-0.7, +4.5) 5 4 (+0.7, +4.5) (-2, +2) Longitudinal* (g) (+2, +2) Lateral (g) (-1, -2) -2-3 (+1, -2) * positive longitudinal quasi-static accelerations are aligned with the direction of flight Figure 4-2. Typical Quasi-Static Design (Maximum Expected) Loads in Flight 4-4 HPD Rev. Initial Release

87 Land Launch User s Guide Section 4 Sinusoidal Equivalent Vibration During Flight The longitudinal and lateral low-frequency sinusoidal vibration environments generated at the spacecraft separation plane during liftoff and flight phases are within the limits defined in Table 4-2. The sinusoidal vibration environment for all major flight events are specifically determined for each mission during the CLA. These results determine the maximum notching in the environment spectra that can be used during spacecraft sinusoidal vibration testing. Table 4-2. Sinusoidal Vibrations at the Spacecraft Interface Frequency Range [Hz] Vehicle Amplitude [g] (Longitudinal and Lateral) Zenit-2SLB Zenit-3SLB Rev. Initial Release HPD

88 Land Launch User s Guide Section Random Vibration Ground Random Vibration for Components Near the Spacecraft Interface The spacecraft is subjected to low frequency random vibrations during transportation by rail at the cosmodrome. Table 4-3 envelopes this random vibration environment. The maximum duration of any rail transfer is six hours. Table 4-3. Random Vibration During Ground Transport When the Spacecraft is Not in the Customer Container Spectral Density of Power [g 2 /Hz] Frequency [Hz] X-X Longitudinal Y-Y Vertical Z-Z Lateral Flight Random Vibration Environment The random vibration environment during flight at the spacecraft interface is enveloped in Table 4-4 and Figure 4-3. Maximum values occur during liftoff and are closely correlated with the acoustic environment. The environment applies to components within 0.5 m (20 inches) from the separation plane along any structural path. This environment is not to be applied to the complete spacecraft as a rigid base excitation. Table 4-4. Random Vibration Environment During Flight Frequency [Hz] Spectral Density [g 2 /Hz] Overall Level 6.8 g rms 4-6 HPD Rev. Initial Release

89 Land Launch User s Guide Section 4 1 0,1 Spectral Power Density [g 2 /Hz] 0,01 0, Frequency [Hz] Figure 4-3. Random Vibration Environment During Flight Rev. Initial Release HPD

90 Land Launch User s Guide Section Acoustics Fairing Volume Average Sound Pressure Levels Maximum acoustic pressures occur during lift off and transonic phases of flight. Acoustic characteristics inside the Land Launch fairings are enveloped in Table 4-5 and Figures 4-4 and 4-5. Table 4-5. Maximum Expected Acoustic Pressure Envelope Inside Land Launch Fairings 1/3 Octave Band Acoustic Pressure Level [db] Center Frequency [Hz] Zenit-2SLB Zenit-3SLB , , , , , , , , , , OASPL Duration 40 seconds 60 seconds Reference: db in respect to 2 x 10-5 Pa (2.9 x 10-9 psi) 4-8 HPD Rev. Initial Release

91 Land Launch User s Guide Section Sound pressure levels, db Note: Overall acoustic pressure level = 140 db 1/3 octave band center frequency, Hz Figure 4-4. Maximum Expected Acoustic Pressure Envelope Inside the Zenit-2SLB Fairing 140 Acoustic pressure levels, db /3 octave band center frequency, Hz Note: Overall acoustic pressure level = 142 db Figure 4-5. Maximum Expected Acoustic Pressure Envelope Inside the Zenit-3SLB Fairing Rev. Initial Release HPD

92 Land Launch User s Guide Section Shock Overview The maximum shock at the spacecraft interface occurs at the moment of spacecraft separation. Other shock inputs, including those associated with fairing jettison and stage separations, are within this envelope. Zenit-3SLB The maximum expected interface shock response spectrums for the 937- mm, 1194-mm and 1666-mm diameter interfaces are presented in Table 4-6 and Figures 4-6a through 4-6c as a function of clamp band tensioning, when using currently available Saab Ericsson Space (Saab) spacecraft adapters. Maximum SC mass and center-of-gravity corresponding to these band tensions are shown in Section 5. The shock environment may differ if other adapters are used. Customers interested in other adapters are encouraged to contact BLS for further information. Table 4-6. Zenit-3SLB Spacecraft Shock Response Spectra (SRS) With Standard SAAB Adapters Shock Response Spectra (g) 937 Interface 1194 Interface 1666 Interface Frequency Band Tension (kn) Band Tension (kn) Band Tension (Hz) kn , , , , , , , , ,800 Q factor of 10 SRS levels are simultaneous in three mutually perpendicular directions As measured at 50 mm (2 inches) from the separation plane on the spacecraft side of the interface 4-10 HPD Rev. Initial Release

93 Land Launch User s Guide Section 4 Figure 4-6a. Zenit-3SLB Spacecraft Shock Response Spectra (SRS) WIth Standard SAAB 937-mm Adapter, Various Band Tensions Figure 4-6b. Zenit-3SLB Spacecraft Shock Response Spectra (SRS) With Standard SAAB 1194-mm Adapter, Various Band Tensions Rev. Initial Release HPD

94 Land Launch User s Guide Section 4 Shock levels at separation plane Response [g] Q= Frequency [Hz] Figure 4-6c. Zenit-3SLB Spacecraft Shock Response Spectra (SRS) With Standard SAAB 1666-mm Adapter, 30 kn Band Tension 4-12 HPD Rev. Initial Release

95 Land Launch User s Guide Section 4 Zenit-2SLB The maximum expected interface shock response spectrums for a single satellite using the Zenit-2 or SAAB 2624-mm interfaces are presented in Table 4-7, with the SAAB information also pictured in Figure 4-6d. The shock environment may differ if other adapters are used, or if more than one satellite is launched at a time. Customers interested in other adapters or group launches are encouraged to contact BLS for further information. Table 4-7. Zenit-2SLB Spacecraft Shock Response Spectra (SRS) Shock Response Spectra (g) Zenit-2 Adapter (Truss) SAAB 2624 Interface Frequency (Hz) SRS Frequency (Hz) SRS Q factor of 10 SRS levels are simultaneous in three mutually perpendicular directions As measured at 50 mm (2 inches) from the separation plane on the spacecraft side of the interface Shock levels at separation plane Response [g] Q= Frequency [Hz] Figure 4-6d. Zenit-2SLB Spacecraft Shock Response Spectra (SRS) With Standard SAAB 2624-mm Adapter Rev. Initial Release HPD

96 Land Launch User s Guide Section Electromagnetic Environment Overview The spacecraft will experience electromagnetic radiation stemming from: The background, or ambient, cosmodrome environment during ground processing; Ground emitters actively used during launch operations on the Zenit launch pad and during launch; The launch vehicle itself Each of these sources is defined below. Coordination It is necessary to coordinate the operation of spacecraft transmitters and other electronic equipment with emissions by the launch vehicle and sources at the cosmodrome. This is performed as part of the integration process. Allowable spacecraft emissions are described in Section 7. Ambient Cosmodrome Electromagnetic Environment The ambient cosmodrome electromagnetic environment varies by location, and changes over time as new equipment is introduced and older equipment is retired. Figures 4-7, 4-8 and 4-9 therefore provide preliminary maximum values for electromagnetic fields levels in Land Launch facilities where the spacecraft will be present: respectively the two available Payload Processing Facilities, the Launcher Assembly Building and the Launch Complex. These environments will be updated during the integration process HPD Rev. Initial Release

97 Land Launch User s Guide Section Field Intensity, db µv/m ,01 0, Frequency, MHz Figure 4-7a. Ambient Electromagnetic Environment within Payload Processing Facility Site Field Intensity, db µv/m ,01 0, Frequency, MHz Figure 4-7b. Ambient Electromagnetic Environment within Payload Processing Facility Area 31 Rev. Initial Release HPD

98 Land Launch User s Guide Section Field Напряженно Intensity, сть электр.по ля, db дбмкв/м µv/m ,01 0, Часто та, МГц Frequency, MHz Figure 4-8. Ambient Electromagnetic Environment within the Launch Vehicle Assembly Building (Area 42) Напряженно сть электр.по ля, дбмкв/м Field Intensity, db µv/m ,01 0, Часто та, МГц Frequency, MHz Figure 4-9. Ambient Electromagnetic Environment at the Zenit Launch Complex (Area 45) 4-16 HPD Rev. Initial Release

99 Land Launch User s Guide Section 4 Launch Vehicle Radio Equipment The Zenit-2SLB has three sets of Sirius telemetry systems which are located on the second stage and operate in a total of five frequencies. The Zenit-3SLB uses the same three sets of Sirius systems on its second stage operating in the same five frequencies, and also uses the BITC-B telemetry system located on the Block DM-SLB third stage that operates in two additional frequencies. Each configuration has a Glonass receiving system, located on the second stage of the Zenit-2SLB and on the third stage of the Zenit-3SLB. Characteristics of these systems are provided below in Tables 4 8, 4-9 and Table 4-8. Characteristics of the Sirius Transmitters (Zenit-2SLB and Zenit-3SLB) Transmitter Characteristic Meter Band Decimeter Band Nominal Frequency (MHz) db Bandwidth (MHz) Modulation Type PCM-FM PCM-FM Max/Min Antenna Gain Coefficient (db) 0/-7 0/-5 Output Power dbw (15 40 W) dbw (10 30 W) Reduced level relative to main signal of spurious and harmonic emissions (db) Table 4-9. Characteristics of BITC-B Telemetry Equipment (Zenit-3SLB Only) Characteristic Transmitter Nominal Frequency (MHz) db Bandwidth (MHz) 1.25 Modulation Type TBD Max/Min Antenna Gain Coefficient (db) TBD Output Power (W) 17 Reduced level relative to main signal of spurious and harmonic emissions (db) 60 Rev. Initial Release HPD

100 Land Launch User s Guide Section 4 Table Characteristics of the Glonass Receiver (Zenit-2SLB and Zenit-3SLB) Characteristic Zenit-2SLB Zenit-3SLB Nominal Frequency (MHz) db Bandwidth (MHz) Receiver Sensitivity at Nominal Frequency (dbw) Max/Min Antenna Gain Coefficient (db) 7/-3 TBD Radio Frequency Environment at the SC Separation Plane The maximum field intensity levels generated by launcher systems at the spacecraft interface plane are provided in Table 4-11, and Figures 4-10 and These account for such factors as the type and orientation of antennas, and the location of the antennas relative to the spacecraft, but do not account for fairing attenuation. The fairings for the Zenit-2SLB and Zenit-3SLB are both aluminum construction, and will attenuate field levels experienced by the SC during pre-launch preparations and after launch until fairing jettison. The degree of attenuation will depend on the size and location of RF windows (mission-specific) and will be analyzed for each mission. Table Maximum Field Intensity Levels Generated by the Launch Vehicle at the Spacecraft Separation Plane, Without Fairing Attenuation Frequency (MHz) Zenit-2SLB Field Intensity (db µ V/m) Zenit-3SLB (TBR) (TBR) All Other HPD Rev. Initial Release

101 Land Launch User s Guide Section 4 Electric field, db µv/m Sirius TM system MHz MHz MHz Sirius TM System 1,010.5 MHz 1,018.0 MHz , ,000 Frequency, MHz Figure Maximum Field Intensity Levels Generated at the Spacecraft Separation Plane by the Zenit-2SLB, Without Fairing Attenuation Electric field, db µv/m Sirius TM system MHz MHz MHz Sirius TM System 1,010.5 MHz 1,018.0 MHz BITC-B TM System 1,026.5 MHz 1,034.5 MHz , ,000 Frequency, MHz Figure Maximum Field Intensity Levels Generated at the Spacecraft Separation Plane by the Zenit-3SLB, Without Fairing Attenuation (BITC-B field intensity levels are TBR) Rev. Initial Release HPD

102 Land Launch User s Guide Section Spacecraft Thermal and Humidity Environments Introduction Described in this section are the thermal and humidity conditions to which the spacecraft will be exposed from arrival at the Baikonur airport through separation on orbit Ground Thermal and Humidity Environments General Overview, Ground Thermal and Humidity Environments Facility Clean Air Systems The spacecraft thermal and humidity environment is actively controlled by facility, transportation and launch pad clean air systems from the time the spacecraft container is offloaded at Baikonur airport through lift off. This supply is also maintained in the case of a launch standby or abort. Table 4-12 provides the temperature and humidity characteristics of each processing milestone or location. Figures 4-12 and 4-13 portray the air conditioning and venting schemes for the Zenit-2SLB payload unit and the Zenit-3SLB ascent unit (shown while integrated with the launch vehicle). The spacecraft is exposed to ambient facility air in the PPF from the time it is unloaded from the shipping container until it is enclosed in the fairing. Land Launch customers may use one of two PPF s (described in Section 6): Site 254 and Area 31. In both locations the temperature, humidity and cleanliness (Class 100,000 or better per FED-STD-209E) of the ambient air is actively maintained by facility clean air systems. The Launcher Assembly Complex (Area 42) also contains a clean room with active temperature, humidity and cleanliness control that is used during mating of the Ascent Unit (or Payload Unit) to the Zenit second stage. Air supply to the fairing is shut off during this mating operation. This clean room is also available for optional customer use in case physical access to the spacecraft is desired through fairing access doors. Transportation Clean Air Systems Clean, conditioned air is supplied to the spacecraft container by a mobile clean air system while being transported between Baikonur airport (Yubileiny) and the PPF. The mobile clean air system is also used to condition the SC enclosure during transport between the PPF and fueling area at Area 31, and to condition the fairing during all moves following payload encapsulation. There may be an interruption in the supply of conditioned air for no more than 60 minutes during loading of the fully assembled launch vehicle onto the transporter/erector inside the Launcher Assembly Complex (Area 42). Cleanliness of the encapsulated environment within the fairing is always maintained at a Class 100,000 or better level per FED-STD-209E HPD Rev. Initial Release

103 Land Launch User s Guide Section 4 Launch Pad Clean Air Systems A clean air system at the launch pad provides conditioned air to the fairing until the erector is lowered and removed at 12 minutes before launch. Conditioned air is maintained from T-12 minutes continuously through launch, and after T-0 in the event of an abort until the transporter/erector and its associated conditioned air system can be reattached to the launch vehicle, by a high pressure payload fairing purge system. A Uninterruptible Power Supply (UPS) system ensures that backup power is available for the pad air supply unit such that conditioned air flow to the fairing can be resumed within one minute after failure of the primary power supply. Impingement Velocity of Airflow Upon SC Surfaces Airflow impingement upon the spacecraft surfaces is generally maintained at or below three meters per second. Rev. Initial Release HPD

104 Land Launch User s Guide Section 4 Table Spacecraft Ground Thermal and Humidity Environment Acting Temperature Relative Operations Phase/Location System [ 0 C] Humidity Nominal Flow Rate SC Container, Airport to PPF Transfer Mobile Unit 15 to 30* 60% 3K-6K m 3 /h Area 31 Area B 15 to 28 35% - 60% PPF Processing Facility Air Area 31 Room to 23 40% - 60% Areas Conditioning Site to 25 30% - 60% N/A Transfer from PPF to HPF (Area 31) Mobile Unit 15 to 30* 60% 3K-6K m 3 /h PPF Fueling Cells Area 31 Facility Air 15 to 25 30% - 60% Site 254 Conditioning TBD TBD N/A PPF Encapsulation Area 31 Facility Air 15 to 28 35% - 60% Halls Site 254 Conditioning 18 to 25 30% - 60% N/A Transfer to ILV Integration Area (Site 42) Mobile Unit 10 to 35* 30% - 60% 3K-6K m 3 /h ILV Integration Bay (Site 42) Facility Air Conditioning 18 to 25 80% N/A Clean Room (Site 42) Facility Air Conditioning 21 to % - 60% N/A ILV ready in Site 42 for roll-out Mobile Unit 10 to 35* < 60% 3K m 3 /h ILV transfer to launch complex (and from launch complex after launch abort) Mobile Unit 10 to 35* < 60% > 2250 kg/h ILV erection, and while erect prior to LOX loading ILV de-erection if launch aborted before T-12 min ILV erect on launch pad, from LOX loading until T- 12 min Zenit-2SLB Zenit-3SLB Zenit-2SLB Zenit-3SLB ILV erect on launch pad, T-12 min to T-0 ILV erect on launch pad following launch aborted between T-12 and T-0, through ILV de-tanking and de-erection (until mobile unit is reconnected) Pad System Pad System High Pressure Pad System 10 to 35* 8 to 25* 10 to 35* 8 to 25* DP C DP C DP C DP C 9500 m 3 /h 5000 m 3 /h 9500 m 3 /h 5000 m 3 /h 10 to 32* DP C > 2250 kg/h Notes: Temperatures maintained within C of set point agreed with the customer * Denotes temperatures as measured at the fairing (or SC container) inlet The customer is responsible for monitoring the environment inside the SC container DP = dew point, K = thousands 4-22 HPD Rev. Initial Release

105 Land Launch User s Guide Section 4 Second Stage Vents Fairing Vents SC Fairing A/C Inlet High Pressure Pad System (from T-12 minutes) Up to 2250 kg/hr Fairing A/C Inlets (to T-12 minutes) Up to 9500 m 3 /hr Separating Screen Second Stage A/C Inlet (Up to 3600 m 3 /hr) Figure Zenit-2SLB Ascent Unit Air-Conditioning (A/C) and Venting Scheme Fairing A/C Inlet (Up to 5000 m 3 /hr) Third Stage A/C Inlet (Up to 4500 m 3 /hr) Second Stage A/C Inlet (Up to 3600 m 3 /hr) Third Stage Vents Second Stage Vents Figure Zenit-3SLB Ascent Unit Air-Conditioning (A/C) and Venting Scheme Rev. Initial Release HPD

106 Land Launch User s Guide Section Flight Thermal Environments General Overview, Flight Thermal Environment After launch, the spacecraft will experience: Heat flux radiated from internal surfaces of the fairing, before fairing jettison. This is mitigated by insulation and ablative coatings on the fairings. After fairing jettison, free molecular heating and various other thermal influences. This is mitigated by the late timing of fairing jettison, by the short duration of the Zenit-2SLB mission, and on Zenit-3SLB by the thermal maneuvering capabilities of the Block DM-SLB Thermal effects experienced by the spacecraft during flight are summarized in Table A thermal analysis will be performed for each mission, using the spacecraft thermal model provided by the customer, to assess spacecraft temperatures during all mission phases. Table Flight Thermal Environments Thermal Effect Zenit-2SLB Zenit-3SLB Thermal flux radiated onto the spacecraft from fairing internal surfaces 500 W/m 2 maximum 400 W/m 2 maximum Free molecular heating at 1135 W/m 2 or less. Considerably less than 1135 W/m 2 for most fairing jettison Free molecular heating after fairing jettison Heat radiated onto spacecraft surfaces by the second stage solid propellant separation thrusters Solar heating, planetreflected solar heating (albedo), Earth-radiated heating, radiation to space missions (typically around 50 W/m 2 ) due to drop zone requirements Typically not significant due to short mission duration. Analyzed for each mission. Dependent on mission profile, but may spike slightly at the time of the second Block DM-SLB burn as shown in Figure Kw-s/m 2 maximum 5.1 Kw-s/m 2 maximum Typically not significant due to short mission duration. Analyzed for each mission. Analyzed for each mission. For thermal management, the Block DM-SLB is designed to accommodate preferred attitude pointing, continuous rolls, maneuvers, and orientations during coast and preseparation phases of flight 4-24 HPD Rev. Initial Release

107 Land Launch User s Guide Section Dispersed free molecular flux, W/m MEO GTO/GSO Time from lift-off, seconds Figure Zenit-3SLB Free Molecular Heating Environment Rev. Initial Release HPD

108 Land Launch User s Guide Section Pressure Venting Overview During ascent, the payload volume is vented through a set of orifices in the second stage equipment bay and in the third stage, as shown in Figures 4-15 and Pressure Decay Rate Pressure Differential at Fairing Jettison The depressurization rate, though varying somewhat by trajectory and dependent on spacecraft displaced volume, does not exceed: kgf/cm 2 per second for Zenit-2SLB kgf/cm 2 per second for Zenit-3SLB A typical fairing cavity pressure curve for the Zenit-2SLB and Zenit- 3SLB are provided in Figures 4-17 and 4-18, along with the associated pressure decay profiles shown in Figures 4-19 and The specific predicted pressure venting rate for each launch is determined during the mission analysis phase. Due to the late timing of fairing jettison due to drop zone constraints, the maximum pressure differential between the pressure inside the fairing and the external pressure at fairing jettison does not exceed a very low kgf/cm 2 for both Zenit-2SLB and Zenit-3SLB HPD Rev. Initial Release

109 Land Launch User s Guide Section 4 Second Stage Vents Separating Screen Figure Zenit-2SLB Ascent Venting Scheme Third Stage Vents Second Stage Vents Figure Zenit-3SLB Ascent Venting Scheme Rev. Initial Release HPD

110 Land Launch User s Guide Section 4 Pressure, kgf/cm Ambient Pressure Ambient Pressure Predicted Internal Pressure Time From Launch, seconds Figure Typical Zenit-2SLB Fairing Internal Pressure Profile During Ascent To be provided Figure Typical Zenit-3SLB Fairing Internal Pressure Profile During Ascent 4-28 HPD Rev. Initial Release

111 Land Launch User s Guide Section 4 Pressure Change Rate, kgf/cm 2 per second Fairing Cavity Pressure, kgf/cm 2 Figure Typical Zenit-2SLB Fairing Internal Pressure Decay Profile To be provided Figure Typical Zenit-3SLB Fairing Internal Pressure Decay Profile Rev. Initial Release HPD

112 Land Launch User s Guide Section Contamination Contamination Control During Ground Processing The spacecraft is protected from contamination during ground processing by: Supplying a continuous flow of clean, conditioned air (class 5,000 or better per FED-STD-209E) to the SC while in its container or under the LV fairings, through launch and after T-0 in the event of a launch scrub or abort. These clean air systems, both mobile and fixed pad units, are described in more detail in Section and maintain a constant overpressure inside the enclosure relative to ambient to prevent outside air ingress. Providing Class 100,000 or better per FED-STD-209E clean room facilities for all spacecraft operations (unloading, processing, fueling, encapsulation) between removal from the SC container and encapsulation in the LV fairing. Precision cleaning of the launch vehicle hardware surfaces that enclose the spacecraft, prior to placing them in proximity to the spacecraft. These cleanliness levels are described in Table Table Fairing Internal Surface Cleanliness Levels at Encapsulation Particle Level 500 Particles Maximum Fairing Surface Levels Level 750 Size per Mil-Std-1246C Zenit-2SLB Zenit-3SLB per Mil-Std-1246C >100 µm 11,900/m 2 11,900/m 2 30,129/m 2 96,300/m 2 >250 µm 281/m 2 281/m 2 753/m 2 2,310/m 2 >500 µm 10.8/m /m 2 32/m /m 2 Zenit-2SLB Zenit-3SLB Non-Volatile Residue 10 mg/m 2 (Level A per Mil-Std-1246C) TBD 4-30 HPD Rev. Initial Release

113 Land Launch User s Guide Section 4 Contamination Control During Flight Potential sources of the SC contamination in flight are fairing materials, contaminants migrating from the launch vehicle equipment bays while venting during ascent, and plume impingement from the second stage solid propellant separation thrusters and, on Zenit-3SLB missions, from the Block DM-SLB steering engines and outgassing. All of these sources are addressed in the design of the Land Launch hardware and mission. Fairing Design Features to Minimize Contamination Materials exposed to the cavity shared by the spacecraft are selected to preclude crumbling, peeling, particle shedding, oxidation or corrosion, and with low outgassing properties that should not exceed the following values during testing according to GOST R (equivalent to ASTM E-595): Total mass loss (TML) less than 1% Collected Volatile Condensable Material (CVCM) less than 0.1% Pyrotechnic devices used for fairing and satellite separation are sealed and do not release gasses or particles. The venting system is designed to preclude circulation from the upper stage (Zenit-3SLB) or second stage avionics bay (Zenit-2SLB) back into the payload unit. Plume Impingement The plume effect on the SC of the second stage solid propellant separation thrusters is negligibly small because of their location on the aft end of the stage. On three stage missions, the Block DM-SLB performs a Contamination and Collision Avoidance Maneuver (CCAM) following satellite separation to ensure a negligibly small contaminating effect on the SC from the upper stage steering engines and stage venting. CCAM features include stage attitude control for optimum orientation relative to the SC, minimum separation distance before main engine re-start, long burns for orbit separation and fuel depletion, and fuel tank venting as a final act. The Block DM family has performed CCAM hundreds of time. Rev. Initial Release HPD

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115 Land Launch User s Guide Section 5 5. SPACECERAFT INTERFACES 5.1 Mechanical Interfaces Overview Mechanical interfaces covered in this section include: mass properties and modal frequencies fairing volumes, access doors and RF windows spacecraft adapters Mass Properties and Modal Frequencies Spacecraft Mass and Longitudinal Center-of Gravity Location Table 5-1 presents the allowable spacecraft mass and longitudinal centerof- gravity (CG) limits for the Zenit-3SLB, based on the standard spacecraft adapters (SCA) currently offered by Land Launch. Customers with spacecraft having a mass / CG configuration that exceeds the defined envelopes should still consult Land Launch for a more detailed assessment. Table 5-2 presents the analogous mass / CG envelope data for the Zenit- 2SLB. SC Mass (kg) Table 5-1. Expected Spacecraft Mass and CG Limits Zenit-3SLB Maximum Spacecraft CG (meters) 937 SCA 1194 SCA 1666 SCA Band Tension (kn) Band Tension (kn) Band Tension kn Notes and assumptions: Mass refers to the separated mass of the spacecraft CG refers to longitudinal distance forward of the separation plane Lateral load factor equal to 2 g Rev. 01/2004 HPD

116 Land Launch User s Guide Section 5 Table 5-2. Expected Spacecraft Mass and CG Limits Zenit-2SLB SC Mass (kg) Maximum Spacecraft CG (meters) 2624-mm SCA Zenit-2 Adapter Notes and assumptions: Mass refers to the separated mass of the spacecraft CG refers to longitudinal distance forward of the separation plane Lateral load factor equal to 2 g. Spacecraft Centerof Gravity Radial Offset The radial offset of the spacecraft CG, relative to the launch vehicle longitudinal centerline, should not exceed: 50-mm (Zenit-2SLB) 25-mm (Zenit-3SLB) Exceptions should be brought to Land Launch for assessment. Modal Frequencies Table 5-3 presents guidelines for spacecraft fundamental natural frequencies on Land Launch. Exceptions should be brought to Land Launch for assessment. Table 5-3. Recommended Spacecraft Fundamental Frequencies Launch System Spacecraft Fundamental Frequencies Longitudinal Lateral Zenit-2SLB > 15 Hz > 5 Hz Zenit-3SLB > 20 Hz > 8 Hz 5-2 HPD Rev. Initial Release

117 Land Launch User s Guide Section Payload Fairing Mechanical Interfaces Payload Fairings Land Launch offers a 4.1 meter (outer diameter) fairing for the threestage Zenit-3SLB vehicle (Figure 5-1) and a 3.9-meter fairing (outer diameter) for the two-stage Zenit-2SLB (Figure 5-2) vehicle. Each fairing has a demonstrated flight record. The general lay-outs of the two fairings are shown in Figures 5-3 and 5-4. RSC Energia Photo Figure 5-1. The Zenit-3SLB Uses the Flight-Proven 17S72 Fairing Made by NPO Lavochkin (shown attached to the Block DM) Rev. 01/2004 HPD

118 Land Launch User s Guide Section 5 SDO Yuznoye Photo Figure 5-2. The Zenit-2SLB Uses a Flight-Proven Fairing Made by PO Yuzhmash 5-4 HPD Rev. Initial Release

119 Land Launch User s Guide Section 5 Longitudinal Separation Joint Fairing Field Joint Spacecraft Adapter Transfer Compartment Figure 5-3. General Lay-out of the 4.1-meter 17S72 Fairing Used on the Zenit-3SLB Rev. 01/2004 HPD

120 Land Launch User s Guide Section 5 Intersection Bay Figure 5-4. General Lay-out of the 3.9-meter Fairing Used on the Zenit-2SLB 5-6 HPD Rev. Initial Release

121 Land Launch User s Guide Section 5 Useable Volume Figures 5-5 and 5-6 present the spacecraft static envelopes for the Zenit- 2SLB and Zenit-3SLB fairings. These envelopes account for: worst-case fairing manufacturing and assembly tolerances in-flight dynamic displacement of the fairing in-flight dynamic displacement of the spacecraft, conservatively assumed to be 50-mm (varies with spacecraft fundamental frequencies and subject to verification during coupled loads analysis) These envelopes should not be exceeded by the maximum dimensions of the spacecraft, including worst case tolerances and expanded thermal blankets, under static conditions (one g longitudinal, zero g lateral). Local excursions outside these envelopes can sometimes be accommodated. Customers are encouraged to contact Land Launch for a specific assessment. Spacecraft interface plane (2624 interface) 830 (Zenit-2 frame) Figure 5-5. Spacecraft Static Envelope within the Zenit-2SLB Fairing (dimensions in millimeters) Rev. 01/2004 HPD

122 Land Launch User s Guide Section 5 SC envelope ø2072 ø (937 interface) 369 (1194 ø3645 interface Transfer Compartment 400 (1666 interface) All dimensions in millimeters ø3679 ø3577 ø3595 ø3723 ø3633 ø3645 Figure 5-6. Spacecraft Static Envelope within the Zenit-3SLB Fairing 5-8 HPD Rev. Initial Release

123 Land Launch User s Guide Section 5 Useable Volume Inside Payload Structure There is room inside the payload support structure (spacecraft adapter and transfer compartment) for satellite motor nozzles. Dimensions will be provided in a future version of this User s Guide. Access Doors Up to two access doors are standard on either Land Launch payload fairing, one door per fairing half. Nominally, the maximum size of each individual door can be: 420-mm by 420-mm on the Zenit-3SLB fairing 500-mm by 500-mm on the Zenit-2SLB fairing Allowable locations for the access doors are shown in figures 5-7 and 5-8. Access doors can typically be used until shortly before the ascent unit (Zenit-3SLB) or payload unit (Zenit-2SLB) leaves the PPF for the launch vehicle assembly building (Area 42), though special arrangements can also be made to open the access doors within the clean room inside Area 42. RF Windows Both Land Launch fairings can accommodate RF windows to enable the testing of spacecraft transmitters on the launch pad. Sizes and locations will be coordinated between the customer and the Land Launch team. Customer Insignia There is external space on either Land Launch fairing for a customer logo or insignia. Details will be coordinated between the customer and the Land Launch team Rev. 01/2004 HPD

124 Land Launch User s Guide Section 5 Location of access doors to SC Location of access doors to SC Location of access doors to SC Fig Zenit-3SLB Payload Fairing Locations for Access Doors 5-10 HPD Rev. Initial Release

125 Land Launch User s Guide Section 5 Развертка Unrolled цилиндрической cylinder of the Zenit-2SLB части головного fairing обтекателя Зона возможного размещения люков доступа Crosshatch = Access door allowable locations 6 I II III IV I I II III IV I Плоскость стыка ГО с переходным отсеком Interface plane between the fairing and the intersection bay Figure 5-8. Zenit-2SLB Payload Fairing Locations for Access Doors (dimensions in millimeters) Spacecraft Adapters Saab Spacecraft Adapters A full range of spacecraft adapters (SCA) made by Saab Ericsson Space is available from Land Launch. Standard offerings are shown in Table 5-4. Other adapters may also be used. Customers interested in other adapters are encouraged to contact Boeing Launch Services. Each standard Land Launch SCA features: a clamp band separation system with a 100% successful demonstrated flight history push-off springs to impart an initial delta-velocity to the spacecraft with respect to the launch vehicle two electrical disconnects at the spacecraft interface Rev. Initial Release HPD

126 Land Launch User s Guide Section 5 Table 5-4. Standard Saab Ericsson Space Spacecraft Adapters Available on Land Launch SCA 937 Mass = 49 kg* SCA 1194 Mass = 58 kg* SCA 1666 Mass = 60 kg* SCA 2624 (Usually Zenit- 2SLB only) Mass = 148 kg * On the Zenit-3SLB, Payload Systems Mass (PSM) also includes the mass of the Transfer Compartment (65 kg) and mission-unique harnessing HPD Rev. Initial Release

127 Land Launch User s Guide Section 5 Zenit-2 Adapter (Frame) for Use with Zenit-2SLB The Zenit-2 frame adapter has been successfully used on a number of missions. The spacecraft is attached to four pyro-locks on a 2062-mm diameter. There are two electrical disconnects at the SC interface. Relative velocity between the separating spacecraft and the launch vehicle is imparted by the simultaneous firing of solid-propellant separation thrusters located on the aft end of the second stage. Multi-Satellite Dispensers for Use with Zenit-2SLB Features of the Zenit-2SLB that make it inherently well suited for launching groups of satellites, as well as secondary payloads, include: Heavy-lift LEO performance (see section 3.5) Large fairing (see section 5.1.2) A structurally robust second stage structure that is designed to support a very large mass (on Sea Launch missions it routinely carries more than 25 tonnes of combined third stage, payload and fairing mass) Land Launch partners PO Yuzhnoye and PO Yuzhmash have extensive experience in designing and building unique, cost-effective payload structures for accommodating and separating multiple payloads. Interested customers are encouraged to contact Boeing Launch Services. Rev. Initial Release HPD

128 Land Launch User s Guide Section Electrical Interfaces Overview Electrical interfaces that are covered in the following sections include: hard-line links (umbilical) radio frequency link in-flight commands, measurements and telemetry electrical power for EGSE bonding and grounding Hard Line Links (Spacecraft Umbilical) Umbilical Circuits Umbilical Use During Processing and Launch Circuits capable of simultaneously supporting the functions defined in Table 5-5 are provided between the spacecraft and hook-up locations for the customer s electrical ground support equipment (EGSE). Function Signal Power (External Spacecraft Power) Power (Battery Charging) Table 5-5. Spacecraft Umbilical Links Circuits Capability or Number & Type Up to 20 twisted shielded pairs, each at 250 ma at 50 V ac or 100 V dc 20 A at 110 V dc 20 A at 110 V dc (from T-12 hours until T-0) 70 A at 110 V dc (all other times) The umbilical circuits defined above are available for customer use at each stop in the launch processing flow as shown in Table 5-6. Place PPF PPF Site 42 Launch Complex Table 5-6. Umbilical Hook-Up Locations and Availability Location for Connecting to Customer-Provided EGSE Processing Cell Processing Cell Control Room Fuel Cell Control Room Encapsulation Hall Clean Room and EGSE Room Bunker (power circuits) Customer Control Room (signals circuits) Notes After attaching the SC to the adapter After attaching the SC and adapter to the upper stage (Zenit-3SLB) While erect on the pad through launch, and after T-0 until lowered in the case of a launch abort 5-14 HPD Rev. Initial Release

129 Land Launch User s Guide Section 5 Umbilical Connectors Umbilical circuits are connected to the spacecraft by separation connectors on the spacecraft adapter. Standard Land Launch spacecraft adapters each accommodate two 61-pin connectors or two 37-pin separation connectors. Other required connectors can be accommodated and will be negotiated on a case-by-case basis. Whichever separation connectors are selected, the customer or spacecraft contractor should procure them and provide the mating halves for the launch vehicle contractor to incorporate into the umbilical and the adapter. Details of umbilical connector interfaces for customer EGSE will be coordinated in the ICD Radio Frequency Links Direct RF communications in C-Band, Ku-Band and/or K-Band between the spacecraft in the fairing on the erect launch vehicle at the pad, and customer EGSE located in the bunker, are provided by: RF window(s) in the fairing A bunker roof antenna, and RF channel equipment linking the bunker roof antenna to connectors in the bunker for customer EGSE The timing of transmissions will be coordinated by Land Launch with the Range authorities. Frequencies, RF window location, and connectors will be coordinated and documented in the ICD In-Flight Commands, Measurements and Telemetry General The launch vehicle will provide the signals and the power to initiate the satellite separation system at the pre-determined point in the mission. No launch vehicle power or command lines will cross the spacecraft separation plane. Separation Verification The launch vehicle will detect spacecraft separation via two diametrically opposed separation switches at the adapter/spacecraft interface. Indication of spacecraft separation will be telemetered via redundant channels. Rev. Initial Release HPD

130 Land Launch User s Guide Section 5 Satellite Environment Measurements Commands Lateral and longitudinal accelerations will be recorded near the interface between the spacecraft and the adapter, telemetered and received. Also recorded, telemetered and received will be information sufficient to determine the acoustics, fairing internal surface temperature, pressure decay, low frequency vibrations, high frequency vibrations and shock environments to which the satellite is exposed during launch. The launch vehicle is capable of issuing up to 2 primary and 2 redundant commands to the spacecraft during flight (in addition to commands for lift-off contact, fairing jettison, readiness for spacecraft separation and spacecraft separation). Command characteristics are as shown in Table 5-7: Table 5-7. Characteristics of Commands from Launch Vehicle to Spacecraft Command Feature Timing of Actuation Accuracy of Actuation Timing Signal Duration Current Voltage Property Any time between lift off and satellite separation + 35 ms ms 0.01 to 1.0 Amperes Up to 34 V 5-16 HPD Rev. Initial Release

131 Land Launch User s Guide Section Electrical Power for EGSE Ground Power Electrical power is provided to the customer at each fixed location in the processing flow, as shown in table 5-8. Table 5-8. Electrical Power Supplies for Customer EGSE PPF Processing Cell PPF Fueling Area Launcher Assembly Bldg Launch Complex Bunker Launch Complex Control Room 220 V +10%/-5% 50 Hz + 2% 380 V+ 2% 50 Hz + 1% 3 Phase 120 V, 20 A, single phase 60 Hz 80 kw 60 kw TBD 80 kw 60 kw TBD 40 kw 40 kw TBD 20 kw 20 kw TBD 20 kw 20 kw TBD Uninterruptible Back-up Power Customer EGSE power is backed-up with a dedicated Uninterruptible Power System (UPS). Characteristics are noted in table 5-9. This system is fully redundant (two UPS units in tandem), each with 15 minutes of battery capability at rated full load. Both 50 Hz and 60 Hz power can be provided, but not at the same time. The UPS supplies three-phase power with a neutral which can be provided to the customer in either single or three phase receptacles. Table 5-9. Uninterruptible Power Supply for Customer EGSE Frequency Characteristics 50 Hz 220/380 VAC, three phase "Y" at 40 KVA with power factor from unity to 80% lagging 60 Hz 120/208 VAC, three phase "Y" at 40 KVA with power factor from unity to 80% lagging Rev. Initial Release HPD

132 Land Launch User s Guide Section Bonding and Grounding Bonding The bonding resistance across the spacecraft to adapter interface is less than 10 milliohms at a current less than 10 milliamps. The resistance between the adapter and the launch vehicle is less than 2 milliohms at a current less than 10 milliamps. Before launch, the launch vehicle is connected to ground with resistance less than 100 milliohms at a current less than 200 milliamps. Grounding At each Land Launch facility where the spacecraft is located or customer EGSE is used during the processing flow, electrically conductive surfaces (threaded studs) are provided for connecting to facility ground HPD Rev. Initial Release

133 Land Launch User s Guide Section 6 6. LAND LAUNCH FACILITIES Overview Land Launch has the advantage of using proven and established facilities at Baikonur Cosmodrome. These include: Krainy Airport for launch personnel arrival and departure Yubileiny Airport for spacecraft and ground support equipment shipment Two payload processing facilities: Site 31 and Site 254. Site 31 will be the primary processing facility until Site 254 facilities improvements are completed. The Block DM-SLB is processed in Site 254. Zenit technical complex located at Site 42 for Zenit processing and mating of the Ascent Unit (Zenit-3SLB) or Payload Unit (PLU) with the Zenit stages followed by check-out of the integrated launch vehicle (ILV). Zenit launch complex located at Site 45 for launching the Zenit-3SLB ILV and the Zenit-2SLB ILV. A map of the Land Launch Baikonur facilities is shown in figure 6-1 N 90 km Yubileinyi airfield Site 31 Payload Processing Site 42 Zenit and ILV Processing Area 254 Payload and Block DM Processing Site 45 Zenit launch complex 75 km Krainy Airport Syrdar-ya River City of Baikonur Legend Railroad Road Figure 6-1. Location of the Principal Land Launch Facilities at Baikonur Cosmodrome Rev. Initial Release HPD

134 Land Launch User s Guide Section Transportation of Personnel and Cargo to and from Baikonur Krainy Airport Personnel fly between Moscow and Baikonur via Krainy Airport (Figure 6-2), which is situated six kilometers to the west of Baikonur city. It can accommodate midsize aircraft for passenger travel throughout the year. Flights are available to and from Vnukovo-1 or Vnukovo-3 airports in Moscow on both commercially scheduled and dedicated charter flights. Land Launch will assist customer personnel in obtaining visas through the Federal Space Agency, and will provide customer representatives with access to the Cosmodrome as well as badges to the required facilities. Figure 6-2. Krainy Airport at Baikonur Yubileiny Airport Yubileiny Airport is located 45 km north of Baikonur city within Baikonur Cosmodrome and is operated by Rosaviacosmos. Its runway, which is 4,500 meters long and 84 meters wide and conforms to International Civil Aviation Organization (ICAO) standards for Class 1 airports, was built to accommodate the landings of the Buran space shuttle. It handles aircraft of all classes for both freight and charter flights, including Boeing 747s and Antonov 124s. Commercial launch customers have used it many times for delivering spacecraft and associated support equipment. The airfield can operate year-round at any time of day. A typical off-load is shown in Figure 6-3. Upon arrival of aircraft, the SC container and associated equipment are offloaded from the aircraft and transferred to railcars that are located approximately 50 to 80 meters from the aircraft. Cranes, forklifts and other necessary equipment are available for these operations. The airport is connected by rail and road to all major cosmodrome facilities. 6-2 HPD Rev. Initial Release

135 Land Launch User s Guide Section 6 Figure 6-3. Spacecraft Off Load at Yubileiny Airport Transportation at the Cosmodrome Rail and road networks connect all Land Launch facilities at Baikonur. Land Launch provides the customer with all necessary transportation of equipment and people on base. Generally, equipment will move between facilities by rail while people will move by road. The spacecraft makes three major moves between facilities: from Yubileiny to the PPF, from the PPF to the launcher assembly building (Area 42) and from Area 42 to the launch complex (Area 45). Spacecraft moves are conducted by rail (Figure 6-4), inside protected enclosures (its own shipping container for the first move, and the fairing for the second and third moves) that are continuously purged with clean, conditioned air as described in Section 4. Rev. Initial Release HPD

136 Land Launch User s Guide Section 6 Figure 6-4. Ascent Unit Transportation with Thermostating Car RSC Energia Photo 6.2 Site 31 Payload Processing Facility Overview Launch Launch s primary Payload Processing Facility (PPF) consists of the existing Site 31 complex of buildings and facilities, which has been used previously to process numerous Western and CIS payloads. Site 254 will become the primary PPF when facility upgrades and improvements are completed. All spacecraft processing, propellant filling operations, pressurization, ordnance preparation, and payload fairing encapsulation operations are conducted here. The PPF has controlled access to ensure compliance with United States governmental security regulations as well as self-imposed customer security requirements and procedures. Major PPF features include: spacecraft processing areas spacecraft fueling area fuel storage room oxidizer storage room control rooms for spacecraft ground support equipment garment change rooms with personnel airlock an encapsulation area office areas for spacecraft personnel The layout of the principal buildings at Area 31 is shown in Figure HPD Rev. Initial Release

137 Land Launch User s Guide Section А Е 40Д 48Б 48 48А 87 Storage Lavochkin NPO Storages 87А Assembly & Test Building (ATB) 51 Support building 40А ATB Annex (Vacuum Chamber) 57 Boiler Facility 40Д ATB Annex (Clean Area) 63 Receiver Facility 40Е Ventilation Facility 87А Uninterruptable Power Supply Facility 43 Charge-Storage Battery Station 105 Transformation Station (6/0.4 kv) 44 Fueling Area 120 IAE Storage 45 Oxidizer storage 122 Refrigerating Center 46 Fuel storage 124 Laboratory Building 48 Cooling Tower 125 Cooling Tower 48А Water Recycling Pump Station 380 Electro-Diesel Station (mobile, 200 kw) 48Б Water Tank 87 Workrooms Figure 6-5. Area 31 Partial Facility Lay-Out Rev. Initial Release HPD

138 Land Launch User s Guide Section 6 Building 40/40D, PPF Building 40D Office areas, Building 44, HPF Buildings 40/40D at Area 31 are used for non-hazardous payload processing. Building 40 has three principal work areas, Area A, B and C, that are shown in Figure 6-6. The SC and equipment are brought into Area C from the airport, unloaded and transitioned into building 40D, room 119 for processing. Rooms 119, 119A and 119B in Building 40D are the usual locations for SC processing and check out prior to fuelling, and are shown in Figure 6-7. After the SC is fueled in Building 44 (see below), it is returned to Building 40, Area A for the beginning of joint operations. In Area A, the SC will be mated to the spacecraft adapter and then to the Block DM in the Zenit- 3SLB configuration. This unit will then be rotated to horizontal and encapsulated. For the Zenit-2SLB configuration, the SC will be mated to the spacecraft adapter and intersection bay, and then rotated to the horizontal position for encapsulation. Air-conditioned office facilities are provided at Site 31. These facilities provide private office and conference space for resident spacecraft personnel teams, including separate office space for the spacecraft manufacturer and satellite customer. International data and voice communications circuits are available. The SC is fueled in the Hazardous Processing Facility, Building 44, located about 300 meters from Building 40/40D. Transfer of the SC back and forth is accomplished inside a conditioned container. A layout of Building 44 is shown in Figure 6-8. Key features of Building 44 include: Clean tent preserving Class 100,000 conditions for the SC Control room with blast-hardened bay window overlooking the fuel island and clean tent Fueling island with spill containment system, hazardous vapor monitors and emergency egress doors Communications, fire-fighting and emergency egress systems Supplies of clean water, liquid nitrogen and facility air Breathing air systems for SCAPE Changing rooms 6-6 HPD Rev. Initial Release

139 Land Launch User s Guide Section 6 - N Building 40 Sanitary inspaction room Building 40А Bridge Crane 50/10 tf (2 ps) Н(hook)= 13.74/14.63 m Clean Rooms Sliding gate 8.4 х 10 (h) Area - A Bridge Crane 5 tf Н=8.1 m Area - С Area - B Sliding gate 5.5 х 7 (h) Building 40Д Figure 6-6. Lay-out of Buildings 40 and 40D

140 Land Launch User s Guide Section 6 Encapsulation Bay Area 300 sq m Height m Class 100,000 clean Crane 1: 5 t, 16.5 m Crane 2: 10 t, 14.6 m Crane 3: 50 t, 13.7 m SC Processing Room Area 240 sq m Height 10.8 m Class 10,000 clean Crane: 5 t, 8.1 m SC Processing Room Area 300 sq m Height m Class 100,000 clean Crane 1: 10 t, 14.6 m Crane 2: 50 t, 13.7 m Figure 6-7. SC Processing and Joint Operations Area in Buildings 40 and 40D

141 Land Launch User s Guide Section 6-1 Filling Hall 3 Oxidizer Loading Area 5 Air Lock 2 Pressurizing Hall 4 Changing Room 6 Shower/Medical Station Figure 6-8. Hazardous Processing Facility, Building 44, at Site 31

142 Land Launch User s Guide Section Site 254 Payload Processing Facility Overview Site 254 will become the primary spacecraft processing facility for Land Launch after various upgrades and improvements have been completed. The upgrades include an additional processing/fuelling cell adjacent to an existing building. A layout of the existing PPF with the proposed processing cell is shown in Figure 6-9. Customer Offices Area 101 Encapsulation Area Area 102 Block DM Processing Loading/Unloading Area Container Cleaning and Acceptance Uninterruptible Power Supply SC Processing/Fueling Area Sumps for Spill/ Run-Off Containment Figure 6-9. Lay-out of SC PPF at Site 254 with proposed adjacent building Site 254 PPF layout The main areas of the PPF for the SC are 101, 102 and the proposed new processing cell. Upon arrival from Yubileini, the SC container and equipment are off-loaded in Area 101 that is located in the central bay of the PPF. Cleaning and acceptance of the cargo is performed in Area 101. The proposed new processing area is located adjacent to Area 102. All SC autonomous operations are performed in this cell. Integrated operations occur in Area 102 as illustrated in Figure Site 254 The PPF is equipped with systems to support all SC processing. Major PPF 6-10 HPD Rev. Initial Release

143 Land Launch User s Guide Section 6 PPF features - technical systems include: Power supply, 380/220 V, 50 Hz; 280/120 V, 60 Hz Compressed gases (air, nitrogen,helium) Conditioned air SC processing area SC fueling area, including remote control room SC storage room Oxidizer storage room Control rooms for spacecraft ground support equipment Garment change rooms with personnel airlock Encapsulation area Office areas for spacecraft personnel Figure Encapsulation Operations in Site 254 Room 102 RSC Energia Photo Rev. Initial Release HPD

144 Land Launch User s Guide Section Zenit Technical Complex Site 42 Overview The Zenit-TM technical complex located within Site 42, which includes the launch vehicle assembly and testing Building 41 (Figure 6-11), is used for: Standalone integration and testing of the Zenit stages Mating of the Zenit with the Ascent Unit (Zenit-3SLB) or with the Payload Unit (Zenit-2SLB), to form the Integrated Launch Vehicle (ILV) Integrated ILV testing ILV loading onto the transporter/erector, prior to moving to the launch complex for launch The complex also includes office space for customer personnel, an equipment room, and a clean room. 3P11107 SDO Yuzhnoye Photo Figure North Rail at the Zenit Technical Complex, Site HPD Rev. Initial Release

145 Land Launch User s Guide Section 6 Integration Area Layout/Features - Building 41 is 120 meters long and 60 meters wide, with three parallel sets of floor-mounted rails. The center rails are used for hardware delivery into and out of the building. The rails on the north side are used for launch vehicle integration operations, while the south rails are currently used for hardware storage. Two traveling bridges each have two cranes, with 50-tonne and 10-tonne capacities. Straddling the north side rail is the clean room (Figure 6-12) that is used for mating the Ascent Unit/PLU to the Zenit second stage. The environmental parameters of this clean room are defined in Section 4. While the fairing is in the clean room the customer has the option of accessing the spacecraft through doors in the fairing. Stands and ladders are available if required. Figure Clean Room at Area 42 SDO Yuzhnoye Photo Spacecraft Equipment Room An equipment room is available for customer use in Building 41, equipped with the power supplies and the umbilical connections to the spacecraft that are defined in Section 5. Rev. Initial Release HPD

146 Land Launch User s Guide Section 6 Customer Office Areas Air-conditioned customer office facilities are provided at in Building 41, Site 42. These facilities provide private office and conference space for resident spacecraft personnel teams. The customer is provided with local and international telephone communication, internal technological communication, broadcasting communication, access to data transmission channels within Baikonur cosmodrome as well as to the international communication channels from Site Zenit Launch Complex (LC) Site 45 Overview Launch Complex Automated Systems A general lay-out of the Zenit launch complex is shown in Figure It consists of two adjacent launch pads supported by shared infrastructure, including propellant tank farms, bunkered launch control complex, and control equipment. Land Launch employs the operational #1 launch pad for both Zenit-3SLB and 2SLB missions. Many features are nearly identical to the ones found on Sea Launch, including launch pad, autocoupling and fueling systems, the transporter/erector and the control system. Launch operations are highly automated on Land Launch just as on Sea Launch. This has many advantages including: short time spent on the pad (approximately 28 hours, unless the customer needs more time for spacecraft testing) inherent safety to personnel, since there is no need to physically approach the launch vehicle high launch-on-time probability If the launch process does experience an anomaly requiring termination, it does so automatically, assuring safety of the launch vehicle, spacecraft, and launch complex. If they are needed, launch vehicle de-fueling operations are also implemented remotely from the control post HPD Rev. Initial Release

147 Land Launch User s Guide Section 6-6 N 10 Land Launch Launch pad #1 2 Launch pad #2 (not in current use) 3 Launch control block (command center) 4 Equipment Bunker 5 Launcher storage bunkers (not in current use) 6 Kerosene storage area 7 Oxidizer storage area 8 Pressure bottle storage 9 Compressor station 10 Air conditioning plant A 7 Rail Connection to Site 42 Figure Lay-out of the Zenit Launch Complex, Area 45 Rev. Initial Release HPD

148 Land Launch User s Guide Section 6 Customer EGSE Room (Bunker) Umbilical connection to the spacecraft (described in Section 5) is provided via the cable mast connected to the Zenit second stage and is disconnected at lift-off of the ILV. RF connection to the spacecraft is made through RF windows in the fairing, and also described in Section 5. The customer EGSE for connecting to these umbilical and RF links is positioned in room 114, an underground equipment room located near the launch pad (Figure 6-14). Room 114 is 10.5 meters by 5.6 meters in size. Though it is in the unmanned area during launch final countdown, it is well protected from the environment generated by the launch. Room 114 Launch Pad Figure Location of Room 114 (Customer EGSE Room) 6-16 HPD Rev. Initial Release

149 Land Launch User s Guide Section 6 Command Center - During pre-launch and launch, SC personnel are located in the Command Center (CC), in rooms 131, 132 and/or 137 as shown in Figure Each of these rooms is more than 60 square meters in size. Customer areas in the CC are equipped with: fire and environmental control systems CCTV monitoring of the launch pad and the ILV connections to spacecraft EGSE in the bunker, for monitoring of SC parameters during the countdown connections to the voice net for customer polling during countdown The CC is two levels down inside a reinforced concrete underground building that provides protection for personnel during launch. Entrance to CC Room for location of Customer personnel and equipment Room 131 Room 132 Room 131 Room 132 Room 137 Room 137 Building 4 Building 4 Figure Customer Location Options in the Launch Command Center Rev. Initial Release HPD

150 Land Launch User s Guide Section Cosmodrome Amenities Visa and Access Authorization Customs Clearance Transportation Consumables Security Schedules Land Launch supports customers in obtaining entry visas to Russia by providing the written invitations. To travel to Baikonur cosmodrome it is necessary to obtain a double or multiple visa. Land Launch also provides customer representatives with access to the Cosmodrome as well as badges to the required facilities. Land Launch supports the customer in obtaining customs clearances at all ports of entry and exit as required for the transport of spacecraft and associated GSE. According to the existing customs regulations, the SC and associated GSE will be brought into Kazakhstan as temporary imports (for re-export) and therefore exempt from duties. Nominal administrative fees may be associated with customs clearance in Russia. If so, such fees are the responsibility of the customer. Any customs or export/licensing processes (export license authorization) in the customer s country of origin for equipment and propellants are the responsibility of the customer. The customer is also responsible for providing all associated packing lists and invoices. All work-related transportation of customer personnel and equipment is provided, starting from arrival at the local airport until departure from the airport. All vehicles for personnel are equipped with air-conditioning and heating systems. If necessary, additional vehicles (e.g., VIP transportation) may be rented in Baikonur. Upon request, and preferably one day in advance, Land Launch can provide transportation to meet atypical customer needs, including night shifts. The customer will be provided in the PPF and/or HPF with de-ionized water, ethyl alcohol, compressed air for tool operation, pressurized nitrogen and helium, breathing air system for SCAPE and clean room garments. The customer should provide his own safety-critical equipment such as SCAPE. Around-the-clock security is ensured to preclude access of unauthorized personnel to the SC. This coverage commences with SC arrival to the Cosmodrome Baikonur airport through launch. Customers are provided with daily and workweek schedules. The typical workweek is six days, Monday through Saturday. Additional working time or other daily/weekly schedules can be arranged on a case-by-case basis HPD Rev. Initial Release

151 Land Launch User s Guide Section 6 External Communications Medical Care Accommodations and Dining - The customer is provided with local and international telephone/facsimile communication, and internet access, and access to allotted commutated ground and satellite international channels to transmit data between Baikonur cosmodrome and the SC customer control center. Usage fees will be coordinated in advance. During the launch campaign, Land Launch provides continuous access to a medical staff that can provide treatment to sick or injured personnel. Land Launch has the capability for an emergency medical evacuation to the United States or Europe if required. The medical center for providing the first treatment is located at Site 254 and at a clinic at Site 2Zh located two kilometers from Site 254. Hotel accommodations are available at the Sputnik hotel ( located in the city of Baikonur, and on base in hotels at Site 2Zh near site 254. The Sputnik Hotel offers 120 comfortable rooms and five suites, one restaurant, a bar, a fitness center, a conference hall, offices, a swimming pool, a sauna, a gymnastic hall, a hairdresser, mountain bikes and a variety of other amenities. The hotels at 2Zh (Figure 6-16) accommodate up to 350 people in comfortable single and double rooms. Site 2Zh also features a café-canteen, a medical clinic, the Baikonur museum and the original buildings used by Yuri Gagarin and Academician Korolev which upon special arrangement can be toured and photographed. Figure Hotel 1 at Site 2Zh Near the Site 254 PPF Rev. Initial Release HPD