russia's dnepr launch vehicle user's guide

Issue 2, November 2001 2

This Users Guide contains technical data, the use of which is mandatory for:

evaluation of spacecraft/Dnepr-1 launch vehicle compatibility; and

preparation of all technical and operational documentation regarding a spacecraftlaunch on Dnepr-1 launch vehicle.

All questions on the issues associated with the operation of the Dnepr Space LaunchSystem that were not addressed in this Users Guide should be sent to the address below:

P.O. Box 7, Moscow, 123022, Russian Federation

Tel.: (+7 095) 745 7258

Fax: (+7 095) 232 3485

E-mail: [email protected]

Current information relating to the Dnepr Space Launch System, activities of InternationalSpace Company Kosmotras, performed and planned launches of Dnepr launch vehicle canbe found on ISC Kosmotras web-site:

http://www.kosmotras.ru

Issue 2, November 2001


Document Change Record

Issue Number Description Date Approved

Issue 2 Dnepr SLS Users Guide, completelyrevised

November 2001 Stanislav I. Us,

Designer General,Dnepr Program


Table of Contents

1. Introduction 102. International Space Company Kosmotras 12

2.1 ISC Kosmotras Permits and Authorities 122.2 Dnepr Program Management. Dnepr Team and Responsibilities 12

3. Purpose, Composition and Principal Characteristics of Dnepr Space Launch System

15

4. Dnepr-1 Launch Vehicle 174.1 General Description 174.2 Spacecraft Injection Accuracy 224.3 Launch Vehicle Axes Definition 224.4 Space Head Module 24

4.4.1 Space Head Module Design 244.4.2 Payload Envelope 26

4.5 Launch Vehicle Flight Reliability 305. Baikonur Cosmodrome 316. Dnepr SLS Ground Infrastructure 35

6.1 Elements and General Diagram of Ground Infrastructure 356.2 SC and SHM Processing Facility 356.3 Spacecraft Fuelling Station 396.4 Launch Complex 40

7. Baikonur Operations Flow 427.1 Preparation of LV for Integration with SHM 437.2 Independent SC Processing 437.3 Preparation of SHM for Mating with SC 447.4 SC / SHM Integration 447.5 Transportation of SHM Containing SC by Transporter Erector to the Silo Launcher

45

7.6 SHM/LV Integration and LV Preparation for Launch 467.7 Pre-launch Operations and LV Launch 47

8. SC/LV Interfaces 49


8.1 Mechanical Interface 498.2 Electrical Interface 51

9. Spacecraft Environments 549.1 Stiffness Criteria (Frequency Requirements) 549.2 Quasi-static and Dynamic Loads 549.3 Vibration Loads 549.4 Shock Loads 579.5 Acoustic Loads 579.6 Temperature and Humidity Conditions and Thermal Effect on Spacecraft

58

9.7 Pressure Underneath LV Fairing 599.8 Gas-dynamic Effect on Spacecraft 599.9 SC/LV Electromagnetic Compatibility 609.10 Spacecraft Tests Required to Meet Dnepr LV Launch Services Requirements

62

10. Ground Qualification Tests 6211. Telemetry and Tracking 6412. Analysis of Flight Results 6513. Range Safety 6614. Transportation of Spacecraft and Associated Equipment. Customs Clearance

69

14.1 Transportation of Dummy Spacecraft to Ukraine for Ground Testing 6914.2 Transportation of Spacecraft and Associated Equipment to Baikonur Cosmodrome

69

15. Principles of Pricing Policy and Contractual Relations 7516. Design and Technical Documents to Be Submitted by Spacecraft Authority 7317. Documents to Be Submitted to ISC Kosmotras by Spacecraft Authority 7518. Launch Campaign Schedule 76


List of Figures

Figure. 1-1 Dnepr-1 Lifts off from Baikonur Cosmodrome. September 26,2000 11Figure 2.2-1 Dnepr Program Operations Diagram 14Figure 4.1-1 Dnepr-1 General View 18Figure 4.1-2 SS-18 1st and 2nd Stages Inside TLC 18Figure 4.1-3 Dnepr-1 Performance Curves for Circular Orbits 19Figure 4.1-4 Mission Profile of Dnepr-1 LV Carrying a Large Spacecraft 20Figure 4.1-5 Mission Profile of Dnepr-1 LV with a Group of Spacecraft 21Figure 4.3-1 Launch Vehicle Major Axes 23Figure 4.4.1-1 SHM Configuration 1 Standard Length 25Figure 4.4.1-2 SHM Configuration 2 Standard Length 25Figure 4.4.1-3 SHM Configuration 1 Extended by 850 mm 25Figure 4.4.1-4 SHM with 2-tier Layout 25Figure 4.4.2-1 Payload Envelope Available within SHM Configuration 1 with

Standard Adapter and GDS27

Figure 4.4.2-2 Payload Envelope Available within SHM Configuration 2 withStandard Adapter and EPM

28

Figure 4.4.2-3 Payload Envelope Available within SHM Configuration 1 Extendedby 850 mm with Standard Adapter and GDS

29

Figure 5-1 Yubileiniy Airfield 31Figure 5-2 Syrdarya River 31Figure 5-3 Cosmodrome Residential Area Town of Baikonur 31Figure 5-4 Kosmonavt Hotel 32Figure 5-5 Baikonur Hotel 32Figure 5-6 Facilities of Sputnik Hotel 32Figure 6.1-1 Dnepr SLS at Baikonur Cosmodrome 35Figure 6.2-1 Layout of AITB, Site 42 36Figure 6.2-2 Layout of AITB, Site 31 37Figure 6.4-1 Dnepr Launch Complex Facilities (Sites 106 and 109) 40Figure 6.4-2 Transporter Erector 41Figure 7-1 Spacecraft Processing and Launch Schedule 42


Figure 7.1-1 Loading of TLC Containing LV 1st and 2nd Stages into SiloLauncher

43

Figure 7.2-1 Off-loading of Spacecraft Container from the Plane 43Figure 7.2-2 Spacecraft Processing 44Figure 7.4-1 SHM Integration 45Figure 7.5-1 SHM Loading into Transporter Erector 45Figure 7.6-1 SHM/LV Integration Inside Silo Launcher 46Figure 7.6-2 SHM Loading into Silo Launcher 47Figure 7.7-1 Dnepr LV Liftoff 48Figure 8.1-1 Typical Design of Spacecraft/Standard Adapter Attachment Point 50Figure 8.2-1 Schematics of SC/SC COE Transit Electrical Links 52Figure 9.7-1 Pressure Change Rate inside Payload Envelope 59Figure 11-1 Tracking Station 64Figure 12-1 Flight Data Processing Center 65Figure 18-1 Launch Campaign Schedule 76


List of Tables

Table 4.1-1 Dnepr-1 Main Characteristics 18Table 4.2-1 Spacecraft Injection Accuracy 22Table 5-1 Typical Routes of Transportation at Baikonur Cosmodrome 33Table 6.2-1 Characteristics of Area A, B and C, AITB at Site 31 38Table 9.2-1 Accelerations at SC/LV Interface during Transportation 55Table 9.2-2 Maximum Quasi-static and Dynamic Accelerations at SC/LV

Interface55

Table 9.3-1 Amplitude of Harmonic Oscillations at SC/LV Interface.Longitudinal Axis (X)

56

Table 9.3-2 Amplitude of Harmonic Oscillations at SC/LV Interface. LateralAxes (Y, Z)

56

Table 9.3-3 Spectral Density of Vibro-accelerations at SC/LV Interface 56Table 9.4-1 Shock Spectrum at Spacecraft Attachment Points 57Table 9.5-1 Acoustic Loads 58Table 9.8-1 Maximum Values of 3rd Stage Motor Plume Parameters Affecting

Spacecraft60

Table 9.9-1 Maximum Residual Levels of RF Emissions 61Table 9.9-2 Maximum Levels of Electromagnetic Emissions 61


Abbreviations

AITB - Assembly, Integration and Test BuildingCOE - Checkout EquipmentECOE - Electrical Checkout EquipmentEPM - Encapsulated Payload ModuleFSUE - Federal State Unitary EnterpriseGDS - Gas-dynamic ShieldGPE - Ground Processing EquipmentGPS - Global Positioning SystemICBM - Intercontinental Ballistic MissileICD - Interface Control DocumentISC - International Space CompanyLCC - Launch Control CenterLSA - Launch Services AgreementLSS - Launch Services SpecificationsLV - Launch VehilceMoD - Ministry of DefenseMoU - Memorandum of UnderstandingNSAU - National Space Agency of UkraineRASA - Russian Aviation and Space AgencySC - SpacecraftSDB - State Design BureauSHM - Space Head ModuleSLS - Space Launch SystemTLC - Transport and Launch CanisterUHV - Ultra High Frequency


1. Introduction

Dnepr Space Launch System (SLS) isdesigned for expedient high accuracyinjection of various single or multiplespacecraft (SC) weighing up to 3.7 metrictons into 300 900 km low earth orbitsinclined 50.5, 64.5, 87.3 or 98 degrees.

The core of the Dnepr SLS is the worldsmost powerful intercontinental ballisticmissile (ICBM) SS-18 or Satan, whichpossesses high performancecharacteristics and mission reliabilityconfirmed through 159 launches (including2 launches under the Dnepr Program).

The SS-18s design and invariance of itscontrol system allowed for creating, on itsbasis, a high-performance Dnepr launchvehicle (LV) equipped with a space headmodule (SHM), which meets modernrequirements for SC injection means.

The actual availability of a considerablenumber of missiles (about 150) that have along operational life, of groundinfrastructure at Baikonur Cosmodromethat comprises processing facilities andlaunch complex, drop zones and grounddata processing complex, as well as of ateam of the companies that developed theDnepr SLS, ensures stable provision oflaunch services.

Dnepr Program implementation isexercised in full compliance with theprovisions of the treaties on reduction andlimitation of the strategic offensive arms(START Treaties).

As a first practical step in the DneprProgram implementation, ISC Kosmotrasjointly with the Russian Ministry ofDefense prepared and launched theBritish Surrey Satellite Technology LimitedUoSAT-12 satellite on a modified SS-18(Dnepr-1) rocket on April 21, 1999.

The following step of the Dnepr Programevolution is mastering cluster launches ofsatellites belonging to different customers.September 26, 2000 marked the Dneprlaunch with five spacecraft: MegSat-1(MegSat s.P.a., Italy), UniSat (LaSapienza University, Rome, Italy),SaudisSat-1A and 1B (Space ResearchInstitute, Saudi Arabia) and TiungSat-1(ATSB, Malaysia).

The Users Guide contains information onbasic characteristics, performance, initialdata for spacecraft integration with thelaunch vehicle as well as spacecraftenvironments within the Dnepr SLS thatwill facilitate the preliminary evaluation offeasibility to utilize the Dnepr SLS for thelaunch of a specific spacecraft.


Figure. 1-1 Dnepr-1 Lifts off from Baikonur Cosmodrome. September 26,2000


2. International Space CompanyKosmotras

2.1 ISC Kosmotras Permits andAuthorities

International Space Company (ISC)Kosmotras (a joint stock company) wasestablished in 1997 by aerospaceagencies of Russia and Ukraine.

ISC Kosmotras performs the developmentand commercial operation of the DneprSLS in accordance with the RussianGovernment Decree # 1156 On Creationof Dnepr Space Launch System datedOctober 5, 1998 and Ukrainian Cabinet ofMinisters decree # 1246 On MeasuresRegarding the Establishment of ISCKosmotras dated November 6, 1997.

The activities on Dnepr SLS wererecognized in the joint statement of theRussian and Ukrainian presidents issuedon May 31, 1997, and on February 27,1998 were incorporated into the Programof Cooperation between Russia andUkraine in the field of research andexploitation of outer space for peacefulpurposes, and into the Russian FederalSpace Program and National SpaceProgram of Ukraine.

In the joint statement of the Russian andUkrainian presidents on cooperation in thefield of rocket and space technology dated12th February 2001, the expansion ofcommercial application of Dnepr launchvehicle was quoted as top priority for longterm cooperation of the two countries inthe field of space related activities.

In the Program of Cooperation betweenthe Russian Aviation and Space Agencyand National Space Agency of Ukraine in

the field of research and exploitation ofouter space for peaceful purposes for2001, which is an integral part of the longterm cooperation program between Russiaand Ukraine for 1998 - 2007, the Programof development and operation of theDnepr Space Launch System was calledas one of the priority projects for 2001.

2.2 Dnepr Program Management. DneprTeam and Responsibilities

The top management body of ISCKosmotras is its Board of Directorscomposed of the members from Russia,Ukraine and the Republic of Kazakhstan -heads of companies incorporated in ISCKosmotras, officials of governmentalentities, as well as leading specialists ofthe program.

Direct Dnepr Program management isexercised by ISC Kosmotras principaloffice located in Moscow, Russia.

Given below are Dnepr team members:

Russia

Russian Aviation and Space Agency(Moscow) - state support andsupervision, provision of facilities andservices at Baikonur Cosmodrome;

Russian Ministry of Defense allocation of SS-18 assets to beconverted into Dnepr-1 launchvehicles, SS-18 storage and Dnepr-1standard launch operations;

Joint Stock Company ASKOND(Moscow) primary entity for DneprProgram management;


Joint Stock Company RosobschemashCorporation (Moscow) - coordination ofSS-18 elimination programs;

FSUE Design Bureau of SpecialMachine Building (St. Petersburg) -primary entity for maintenance oflaunch complex and processingfacilities;

FSUE Central Scientific and ResearchInstitute of Machine Building (Moscow)- scientific and technical support of theprogram;

FSUE Scientific and ProductionAssociation IMPULSE (St.Petersburg) - development andupgrade of launch control and supportequipment;

State Enterprise Moscow Electrical andMechanical Equipment Plant (Moscow)- modification of launch vehicle controlsystem instrumentation;

Ukraine

National Space Agency of Ukraine(Kiev) - state support and supervision;

State Design Bureau (SDB) Yuzhnoye(Dnepropetrovsk) - primary design anddevelopment organization for thelaunch vehicle and the entire DneprSLS;

State Enterprise ProductionAssociation Yuzniy Machine BuildingPlant (Dnepropetrovsk) - primarymanufacturing entity;

Scientific and Production EnterpriseKHARTRON-ARKOS (Kharkov) -primary entity for launch vehicle controlsystem;

Republic of Kazakhstan

Aerospace Committee of the Ministryof Energy and Mineral Resources -state support and supervision;

State Enterprise INFRAKOS(Baikonur) - participation in DneprProgram activities at BaikonurCosmodrome;

State Enterprise INFRAKOS-EKOS, asubsidiary of INFRAKOS (Alma-Aty) ecological support of the program.

Additionally, ISC Kosmotras, on a contractbasis, uses services of partners locatedoutside of Russia and Ukraine for DneprProgram marketing.

Figure 2.2-1 shows the Dnepr Programoperations structure. ISC Kosmotrasprincipal office maintains contacts withlaunch customers, conducts preliminaryevaluation of launch services provisionopportunities, concludes contracts withcustomers, organizes their fulfillment bycompanies subcontractors from Russia,Ukraine and Kazakhstan, works incooperation with the Russian Ministry ofDefense, Russian Aviation and SpaceAgency, National Space Agency of theUkraine and National AerospaceCommittee of the Republic of Kazakhstan,as well as with other governmentalagencies of these countries.

In August 1999 ISC Kosmotrasestablished an ISC Kosmotras subsidiarythat is located in the city of Kiev, Ukraine.It is licensed by the NSAU to conductspace related activities. Its primarymission is to participate in the process ofestablishing favorable conditions forpromotion of Dnepr launch services to themarket.


Figure 2.2-1 Dnepr Program Operations Diagram


3. Purpose, Composition and PrincipalCharacteristics of Dnepr Space LaunchSystem

Dnepr SLS is designed for expedient,high-accuracy injection of various single ormultiple spacecraft weighing up to 3.7metric tons into low-earth orbits inclined50.5, 64.5, 87.3 or 98 degrees.

Dnepr SLS is developed on the basis ofthe SS-18 missile system, which is beingdecommissioned in the process ofreduction and elimination of strategicoffensive arms. In the course of thedevelopment of the system, the followingmajor principles are being followed:

maximum heritage with previouslydeveloped and proven systems andunits;

utilization of proven work technologies;and

utilization of LV standard flighttrajectories and their associated dropzones.

Dnepr SLS consists of the followingelements:

launch vehicle with the space headmodule;

launch complex with the launch controlcenter;

launch vehicle, spacecraft and spacehead module processing facilities; and

set of data collection and processingmeans.

Dnepr LV is based on the SS-18 ICBMmodified in order to ensure optimalspacecraft integration and injection withthe minimum costs incurred.

Launch complex is a combination oftechnologically and functionallyinterrelated systems, components,facilities and service lines required tomaintain the readiness status of thelaunch vehicle, and to prepare andexecute its launch.

LV processing facility incorporates specialstructures and mobile processingequipment required to prepare the LV forlaunch.

SC and SHM processing facility isdesigned to perform the followingoperations:

acceptance, temporary storage andprocessing of the spacecraft; and

assembly and processing of the spacehead module consisting of thespacecraft, adapter, intermediatesection, gas-dynamic shield and LVfairing.

Available telemetry posts are used forreceiving telemetry data during injectioninto orbit and for taking trajectorymeasurements.

If the spacecraft is launched southward(87.3, 98 degrees inclination), telemetryposts located along the flight path on theterritory of foreign countries may be usedin order to ensure stable coverage.

The sequence of ground telemetry postoperation (including mobile telemetryassets) is contingent on the injectionpattern of a specific spacecraft.


Depending on the spacecraft launchprogram requirements 1 to 3 silolaunchers can be used, which are capableof up to 25 launches per year (if thepersonnel work in one shift).

The launch vehicle inside its launchcanister, when fuelled and placed insidethe silo launcher, can be on stand-byawaiting integration with the SHM and SCfor an unlimited period of time throughoutits operational life.


4. Dnepr-1 Launch Vehicle

4.1 General Description

Dnepr-1 launch vehicle with a space headmodule is a basic modification of theliquid-fuelled SS-18 ICBM with a three-stage-plus-SHM in-line configuration.

General view of the launch vehicle with thespace head module is shown in Figures4.1-1 and 4.1-2.

The LV first and second stages arestandard SS-18 elements and usedwithout modification.

First stage propulsion unit features foursingle-chamber motors, while the secondstage propulsion unit is composed of amain single-chamber motor and a four-chamber thruster.

The LV third stage is a modified standardSS-18 third stage equipped with a liquidpropellant, two-mode propulsion unit thatoperates based on a drag scheme.Modifications involve only the controlsystem in order to provide optimal flightsoftware and electrical links with thespacecraft.

SHM is attached to the third stage upperend. SHM consists of a spacecraft,intermediate section, adapter, either gas-dynamic shield (GDS), or EncapsulatedPayload Module (EPM), protectivemembrane and SS-18s standard fairing.

Dnepr LV features a standard inertial high-precision computer-based control system.

The LV telemetry system ensurestransmission of telemetric data from theLV up to SC separation (including themoment of separation) from the LV.

The LV safety system ensures flight abortof the first and second stages in case ofan emergency (loss of flight stability). Thissystem is based on the safety systemused for flight testing of the SS-18 ICBM.

Dnepr LV is steam ejected from itstransport and launch canister to a height ofapproximately 20 meters above theground by means of activation of the blackpowder gas generator. The first stagepropulsion unit is ignited upon the rocketejection from the launch canister.

Separation of stages and fairing followsthe proven SS-18 procedures. Spacecraftseparation from the third stage is done bythe third stage taking away from thespacecraft by means of throttled-backoperation of its motor. Prior to thespacecraft separation, the gas dynamicshield or EPM cover is jettisoned.

Principal characteristics of Dnepr-1 LVwith a single-tier space head module aregiven in Table 4.1-1.

Dnepr-1 LV performance curves arepresented in Figure 4.1.3.


Figure 4.1-1 Dnepr-1 General View

Table 4.1-1 Dnepr-1 Main Characteristics

Liftoff mass (with the spacecraft mass of 2000 kg), kg

1st stage 208900

2nd stage 47380

3rd stage 6266

Thrust in vacuum, tons

1st stage 461.2

2nd stage 77.5

3rd stage (primary mode/throttledback operation mode)

1.9/0.8

Propellant components for all stages

Oxidizer Amyl

Fuel Heptyl

Effective propellant capacity, kg

1st stage 147900

2nd stage 36740

3rd stage 1910

Flight reliability 0.97

SC injection accuracy (Orbit altitude Hcir =300 km)

for orbit altitude , km 4.0period of revolution, sec. + 3.0

for inclination, degrees 0.04for ascending node right ascension,degrees

0.05

Orbit Inclination, degrees 50.5; 64.5;87.3; 98

Figure 4.1-2 SS-18 1st and 2nd StagesInside TLC


0

400

800

1200

1600

2000

2400

2800

3200

3600

4000

300 400 500 600 700 800 900

Figure 4.1-3 Dnepr-1 Performance Curves for Circular Orbits

Gsc, kg

cir, km

i=50.5 i=64.5 i=87.3 i=98.0


Figure 4.1-4 Mission Profile of Dnepr-1 LV Carrying a Large Spacecraft


Figure 4.1-5 Mission Profile of Dnepr-1 LV with a Group of Spacecraft


4.2 Spacecraft Injection Accuracy

In spacecraft mission profile to a circularorbit without application of yaw maneuver,the maximum deviations (with probabilityP=0.993) of orbital parameters fromnominal values at the moment ofspacecraft separation do not exceed thevalues given in Table 4.2-1 below.

In case of multiple spacecraft injection, theorbit is calculated for each spacecraft, andthe injection accuracy of a specificspacecraft may be specified.

The data contained herein is of generalcharacter. For each specific flight, thecomposition and values of monitoredparameters may be specified subject tolaunch mission and associatedmodifications of 3rd stage control systemoperation pattern, required orbitparameters, payload mass (number ofspacecraft being inserted), separatedelement drop zones, etc.

4.3 Launch Vehicle Axes Definition

The Launch Vehicle pitch and yaw planesas well as its rotation axes are defined bythe LV center line as its longitudinal axisand by mutually perpendicular I-III and II-IV axes (see Figure 4.3.1):

the pitch plane is defined by the LVlongitudinal axis and I-III axis.

I-III axis is positioned in such a way thatthe I-direction of the axis is pointing awayfrom the launch site (lift-off point), whereasIII-direction is pointing to the launch site;

the yaw plane is defined by the LVlongitudinal axis and II-IV axis.

The LV angular displacements aredetermined (relative to an observer lookingforward from the LV aft section) as follows:

the LV makes pitch rotation around II-IV axis. The pitch positive movementdisplaces the LV nose section in theupward direction (towards direction IIIof I-III axis);

the LV makes yaw rotation around I-IIIaxis. The yaw positive movementdisplaces the LV nose section in theleft-hand direction (towards direction IIof II-IV axis);

the LV makes roll rotation around thelongitudinal axis. The roll positivemovement turns the LV clockwise(from direction I of I-III axis to directionII of II-IV axis).

Positive directions of the pitch, yaw androll rotation are shown by arrows in Fig.4-3.1.

Table 4.2-1 Spacecraft Injection Accuracy

Typical circular orbitsOrbital Parameter

H=300 km, i=98o H=600 km, i=98o H=900 km, i=65o

Altitude, km 4.0 5.5 10.0Period of revolution, sec. 3.0 4.0 6.5Inclination, deg. 0.040 0.045 0.050Right ascension of the ascending node, deg. 0.050 0.060 0.070


Figure 4.3-1 Launch Vehicle Major Axes

Positive PitchDirection

Positive YawDirection

Positive RollDirection


4.4 Space Head Module

4.4.1 Space Head Module Design

The spacecraft is installed inside the SHM.The SHM is composed of the fairing,cylindrical intermediate section, adapter,protective membrane, GDS or EPM.

Fairing is a four-cone structure that has alongitudinal joint along stabilization axes Iand III, which divides the fairing into twohalf-shells (sections) that are tied togetherby 28 pyro-devices. The fairing is installedatop the cylindrical intermediate sectionand attached to it by means of 8 pyro-devices. Upon the fairing separationcommand, the half-shell andfairing/intermediate section attachmentpyros are activated, the half-shells arehinged by means of spring pushersinstalled at the bottom end of the fairingand, upon reaching a certain angle of turn,are separated from the intermediatesection.

Intermediate section is a cylindrical partthat has a length of 2080 mm (standardsize) and diameter of 3000 mm andincorporates two platforms A (on upperextreme ring frame of which, the fairing isinstalled) and B (the bottom extreme ringframe of which is attached to the 3rdstage). Both platforms are interconnectedby 6 pyro-devices. The length ofintermediate section can be extended by850 2000 mm to incorporate a biggerspacecraft.

Adapter is a newly developed conicalstructure. Attached to the adapter upperflange are the spacecraft and a specialprotective membrane that isolates thepayload envelope from the control andtelemetry system instrumentation

compartment located under the adapter.The adapter bottom flange sits on thelower extreme ring frame of the platform B.

In necessary, an additional adapter (orseveral adapters) that is already integratedwith the spacecraft, may be placed on thestandard adapter.

SHM is available in two configurations:

Configuration 1 a newly developedintermediate section with GDS;

Configuration 2 an intermediate sectionconsisting of standard platforms A and Bplus EPM.

Both configurations of SHM use thestandard fairing and a newly developedadapter.

For SC protection against the 3rd stagemotor plume, the Configuration 1 utilizesthe GDS that is attached to the Platform Aupper ring frame and separated prior tothe SC release.

For the above purpose, the Configuration2 uses the EPM, the cover of which isseparated prior to the spacecraft release,i.e. similarly to the GDS.

Layout schematics of the standard lengthSHM (both with GDS, and EPM) is shownin Figure 4.4.1.1 and 4.4.1.2 respectively.

Layout schematic of SHM Configuration 1,the length of which is extended by 850mm, is shown in Figure 4.4.1.3.

SHM design allows for multi-tier spacecraftlayout. One of the options for such layoutis shown in Figure 4.4.1.4.


Figure 4.4.1-1 SHM Configuration 1 Standard Length

Figure 4.4.1-2 SHM Configuration 2 Standard Length

Figure 4.4.1-3 SHM Configuration 1 Extended by 850 mm

Figure 4.4.1-4 SHM with 2-tier Layout


4.4.2 Payload Envelope

Payload envelope is a volume within theSHM, which is designed foraccommodation of spacecraft.

Spacecraft dimensions (including all of itsprotruding elements) must fit within thespecified payload envelope, given allpossible deviations and displacementsfrom the nominal position during theground testing and in flight.

The size of payload envelope within thestandard SHM (length 5250 mm) and

adapter (H = 550 mm) is shown in Figures4.4.2.1 and 4.4.2.2 (i.e. for SHMConfiguration 1 with GDS and SHMConfiguration 2 with EPM respectively).

Figure 4.4.2.3 shows the size of payloadenvelope within the SHM extended by 850mm (maximum possible extension of theSHM length is 2000 mm).

When designing the interface between thelaunch vehicle and a specific spacecraft,the size and configuration of the payloadenvelope may be specified.


Figure 4.4.2-1 Payload Envelope Available within SHM Configuration 1 with StandardAdapter and GDS


Figure 4.4.2-2 Payload Envelope Available within SHM Configuration 2 with StandardAdapter and EPM


Figure 4.4.2-3 Payload Envelope Available within SHM Configuration 1 Extended by850 mm with Standard Adapter and GDS


4.5 Launch Vehicle Flight Reliability

SS-18 in its basic configuration has beenin operation since mid-70s. Throughoutthe entire period of its operation, a numberof measures have been taken to maintainthe technical condition of the rockets,launch and processing equipment. Thesemeasures include the following:

Regular inspections of the operabilityof the systems, units and equipment;

Scheduled (routine) maintenance;

Replacement of faulty units, modulesand components; and

Periodical checks of the readiness ofall systems, units and equipment forthe performance of their basicfunctions.

The effectiveness and adequacy of theabove measures have been demonstratedby the successful launches of the rocketswith the different storage periods atvarious times during the SS-18 operationallife.

In order to monitor the actual degree of thelaunch vehicle reliability, the leadinginstitutes of the Russian Ministry ofDefense and Russian Aviation and SpaceAgency, as well as State Design BureauYuzhnoye as the principal developer of thesystem, conduct annual independentevaluation of the reliability data values, oftheir compliance with the establishedrequirements and also develop programsof work necessary to ensure the required

level of reliability. The implementation ofthis work ensures and allows to maintainthe high level of the rocket reliabilitythroughout the entire period of itsoperation. Currently, the mission reliabilityfactor for the rocket is 0.97, which hasbeen established through a great numberof successful launches.

The total of 159 launches of the SS-18were carried out as of October 2001 withonly 4 of the launches encounteringmalfunctions or anomalies of certainsystems. The causes of the failures havebeen unequivocally established. Theywere due to isolated manufacturingdefects. Based on the analysis of thesedefects, certain measures were taken toenhance rocket fabrication quality andensure full compliance with engineeringdocumentation during fabrication. Theeffectiveness of the measures taken wasproved through subsequent successfullaunches. No similar recurringmalfunctions were encountered. 8 SS-18launches have been conducted from 1990to 2000, all of them successfullyaccomplished their mission.

The conversion of the SS-18 into Dnepr-1launch vehicle requires minor designmodifications associated with thespacecraft integration with the launchvehicle and spacecraft separation process.The modified components undergovigorous ground testing that enables tomaintain the achieved mission reliabilitylevel. The above-mentioned 8 launchesincluded two launches in Dnepr-1configuration that were performed in 1999 2000.


5. Baikonur Cosmodrome

Dnepr SLS operates from BaikonurCosmodrome, one of the largestspaceports in the world.

Baikonur Cosmodrome is located inKazakhstan east of the Aral Sea (63 E and46 N) in the semiarid zone with sharplycontinental climate. Typical for this areaare hot dry summers and frosty winterswith strong winds and little precipitation. Insummer, the temperature can rise up toplus 450C and in winter, it can drop tominus 400C. Yearly average temperatureis approximately plus 130C.

The Cosmodrome is located sufficiently faraway from the large centers of population.This fact ensures safety of the launchesand facilitates the task of creating bufferareas, which are used as launch vehiclestage drop zones. Another advantage ofthe Cosmodrome location is the fact thattypical for this location is a great deal ofclear days on a yearly basis.

The residential area of the Cosmodrome isthe town of Baikonur located on the rightbank of the Syrdarya River. Directlyadjacent to the town are the village ofTyuratam and the railway station of thesame name. Two Baikonur airports areconnected with the city of Moscow byregular flights.

Figure 5-1 Yubileiniy Airfield

Figure 5-2 Syrdarya River

Figure 5-3 Cosmodrome Residential Area Town of Baikonur

These airports are equipped toaccommodate any types of cargo andpassenger planes.

The town has the following facilities: amovie theatre, a stadium, a swimming


pool, TV center, medical facilities, cafesand restaurants as well as internationaltelephone, telex and facsimile services.Baikonur visitors can stay at several hotelsthat offer single and double rooms.

Accommodations are offered forspacecraft processing specialists andlaunch guests at Sputnik, Baikonur andKosmonavt hotels located in the northernpart of the town.

All Russian and foreign cosmonauts stayat Kosmonavt hotel during preparation forspace flight.

The rooms have air conditioning, atelevision set, a refrigerator and a bathroom. Deluxe accommodations, in addition

Figure 5-4 Kosmonavt Hotel

Figure 5-5 Baikonur Hotel

to the above-mentioned facilities, have abed room and a private office. The cost ofaccommodation is $30 80 per day atBaikonur and Kosmonavt hotels and $220 250 per day at Sputnik hotel.

Figure 5-6 Facilities of Sputnik Hotel


ATMs and exchange offices are availablein Baikonur for currency exchange andcash withdrawal. Both Russian andKazakhstan currencies are accepted. Mostpayments are made by cash.

The Cosmodrome is connected with othercities and countries by air, rail and roadtransportation. There is an extensivenetwork of motor and railroads on theterritory of the Cosmodrome.

The town provides centralized water (hotand cold) and 220V/50Hz power supply.

The basic accommodation for thecustomers personnel is assumed to beKosmonavt hotel.

Table 5.1 shows typical routes oftransportation at Baikonur Cosmodrome.

Table 5-1 Typical Routes of Transportationat Baikonur Cosmodrome

Krainiy Airport Kosmonavt Hotel 6 km

Yubileiniy Airport Kosmonavt Hotel 55 km

Tyuratam RailwayStation Kosmonavt Hotel

3 km

Kosmonavt Hotel Launch Site 45 km

Kosmonavt Hotel Facility 31,SC Fuelling Facility

60 km

Kosmonavt Hotel Facility 42, AITB 55 km

Minivans and sedan cars are available fortransportation of personnel. The presenceof local escorts to accompany transportvehicles is a must.

Accommodation for VIPs is available at 5-star Sputnik hotel. Police escorts can be

arranged for VIP transportation.

During spacecraft preparation for launch atS/C processing facility and launchcomplex, the Customer can be providedwith the following telecommunicationservices:

local/international telephone/facsimilecommunication;

internal technological and publicaddress system between work sites atprocessing facility premises (intercom);

mobile radio and radiotelephonecommunication;

communication channels to transmitUTC and pre-launch countdownsignals;

access to the Internet, etc.

Dnepr SLS telecommunication system isintegrated into the single distributedinformation network of the BaikonurCosmodrome based on the ISDNprinciples by wire cable and radio relaycommunication channels network. Thisnetwork covers the premises and facilitiesof the launch complex (Sites 103, 106,109, 111/2), S/C processing facility (Sites31, 42) and other Cosmodrome facilities.

International communications are providedby a satellite segment based on theIntelsat and Inmarsat type ground stationswith direct access to the Moscowtelephone network.

The mobile communication is provided bythe trunking radio stations covering thecomplete Cosmodrome area, the GSM-900 cellular communication is provided atthe most of the cosmodrome and adjacenttown areas.


A fully equipped military hospital is locatednear the Kosmonavt hotel. Its personnelhas the required qualification andexperience. First medical aid is availableat work premises and sites.

Upon customers request, medicalpersonnel and equipment can be placedon stand-by at spacecraft processingfacility and in emergency circumstances,the patient may be evacuated to aEuropean medical facility.


6. Dnepr SLS Ground Infrastructure

6.1 Elements and General Diagram ofGround Infrastructure

The Dnepr SLS ground infrastructureincludes the following facilities:

SC and SHM Processing Facilities;

SC Fuelling Station;

Launch complex;

LV Processing Facility;

Fueling station for storage, preparationand discharge of rocket propellantcomponents;

Figure 6.1-1 shows infrastructure facilitiesat Baikonur Cosmodrome, which are usedby ISC Kosmotras for Dnepr launchcampaigns.

6.2 SC and SHM Processing Facility

SC and SHM processing facility isdesigned to perform the followingoperations:

Receiving, temporary storage andprocessing of the spacecraft, including,if necessary, its filling with compressedgases and propellant components; and

Assembly and processing of the spacehead module (consisting of spacecraft,adapter and launch vehicle fairing).

94

93

97 90

91

92 82

95

81131

182175

200

251251

250

110

112A118

119254

113

11251

110

1812

104

106, 109

103

45

42 44

43 41

5

3G3K

3R

2321

10

15

111/2

TyuratamBAIKONUR

31 32

Dnepr LV Silo Launchers

Dnepr LV Control and Observation Post

Yubileiny Airfield

Krayniy Airfield

SHM and SC Processing Facility

Kosmonavt,Baikonur,SputnikHotels

YuzhnayaHotel

Dyurmentyube

RailwaysFreewayMotor roads

N

S

EW

Figure 6.1-1 Dnepr SLS at Baikonur Cosmodrome


Containers with the spacecraft andspacecraft ground processing equipment(GPE) can arrive Baikonur via two airports- Yubileiniy and Krainiy as well as viaTyuraTam railway station.

If a spacecraft needs to be kept in certaintemperature conditions, it will arrive atYubileiniy airport, from where it will betransported to SC and SHM processingfacilities by special environmentallycontrolled railcars for subsequentoperations.

If there is no strict temperaturerequirement, the spacecraft and SC GPEwill be transported by road vehicles orgeneral purpose railcars.

Two airports (Yubileiniy and Krainiy) arecapable of receiving all passenger andcargo airplanes, including heavy ones (An-

124 and Boeing 747) and possess allnecessary ground support equipment forhandling operations.

Directly adjacent to Yubileiniy airport is arail line that can be used for transportationof spacecraft and SC GPE to SC and SHMprocessing facility.

Depending on spacecraft processingrequirements, its weight and dimensionalcharacteristics and the cost of operation,ISC Kosmotras offers customersassembly, integration and test buildings(AITB) for spacecraft processing at Site 42and/or Site 31.

If necessary, other AITBs available atBaikonur Cosmodrome may be used forspacecraft processing.

Premises and equipment located inside

Figure 6.2-1 Layout of AITB, Site 42


the Zenith LV AITB were used forspacecraft processing in the course of firsttwo Dnepr launch campaigns (Site 42). Alayout of AITB, Site 42 is shown in Figure6.2-1. Technical data for clean roomfacility (primary place for spacecraftprocessing and assembly of EPM) andother AITB premises is given below.

Clean room facility located inside the AITBis divided into two clean rooms with theaggregate floor area of 212 sq. meters andan airlock compartment of 32.6 sq. meters.Telphers with lifting capacity of 1000 kgfare available in the airlock compartmentand the two clean rooms. Maximum heightto the telpher hook is 8.2 meters. Thesizes of the airlock compartment gate, thegate between clean rooms and the gateopposite the airlock compartment are5.2x6.3 m, 5.2x8.2 m and 5.2x6.3 mrespectively. The temperature inside theclean room facility is maintained within therange of 21.1C - 26.7C with the relativehumidity being 30-60%. Cleanliness classis 100,000 (US FED-STD-209E). Clean

room facility has an anti-static floor, a low-impedance grounding contour (up to 4ohms) and 12V/220V AC 50 Hz electricoutlets. Illumination of airlock compartmentis 300 lux and of clean rooms is 500 lux.

Spacecraft checkout equipment (COE)room has the aggregate floor area of 53sq. meters (29 sq. meters and 24 sq.meters).

The AITB has office premises forpersonnel involved in spacecraftprocessing and assembly (4 rooms, 21 m2each), which are equipped with requiredcommunications means.

The AITB is equipped with the followingutilities: uninterrupted power supply (withthe same characteristics as both in Russiaand the US), heating, air conditioning andventilation, water supply, sewer, securityand fire alarms, and variouscommunication systems.

Near the clean room facility, there is a

Figure 6.2-2 Layout of AITB, Site 31


place where the SHM or EPM isassembled and loaded into the transportererector.

If the weight and dimensions of the ofspacecraft do not allow to use the AITB atsite 42 or the spacecraft filling withpropellant is required during itsprocessing, ISC Kosmotras offers to usethe AITB and fuelling station located atSite 32, shown in Figure 6.2-2.

This AITB consists of three areas A, B andC.

Area A is designed for integration ofEPM and for electrical checks ofspacecraft as part of the EPM. Tomaintain the required cleanlinessconditions during operations, the AreaA is equipped with air shower;

Area B is designed for the spacecraftprocessing and assembly of SHM;

Area C is designed for receiving andtemporary storage of shippingcontainers with SC and SC GPE, ofSHM platforms (SC adapters) andlaunch vehicle fairing, as well as for

handling operations including theloading of the integrated SHM into thetransporter erector for transportation tothe silo launcher.

Main characteristics of areas A, B and Care given in Table 6.2-1 below.

Facility 40D within the Area B consists ofhall 119, 119A and 119B, the area ofwhich is 240, 100 and 100 m2 respectively.Hall 119 will incorporate the spacecraft,hall 119A will host spacecraft COE, andhall 119B will be for mechanicalequipment, devices and consumables.

Hall 119 is equipped with compressed gasfilling system, including filling withNitrogen. This Nitrogen will be suppliedunder the pressure of 40 MPa (standard 92-1577-78) and have the followingparameters:

nitrogen content no less than99.95%;

oxygen content no more than 0.05%;

water vapor 0.004%.

Table 6.2-1 Characteristics of Area A, B and C, AITB at Site 31

Area A B C

Characteristics

Dimensions Length 56.0 mWidth 11.5 mLength 56.0 mWidth 18.5 m

Length 63.0 mWidth 30.0 m

Cleanliness Class (USstandard 209E)

100,000,if necessary 30,000

100,000,if necessary 30,000

-

Height to Crane Hook 13.8 m 13.8 13.8

Crane Lifting Capacity 10 and 50 tons 10 and 50 tons 10 and 50 tons

Temperature 18-25C 18-25C 18-25C

Relative Humidity at 20 C 30-60% 30-60% -


Hall 119 has the following power supplysystem:

120 V 20 A 50 Hz;

208 V 30 A 50 Hz;

15 kWt 50 Hz.

Hall 119 is equipped with video cameras.

The personnel directly involved inspacecraft processing operations at Halls119, 119A and 119B enter area B throughair showers located at both ends of theAITB and separating clean and notclean areas. The personnel notcompletely involved in operations with thespacecraft, who operate communicationsmeans or computers are located in officepremises of the not clean area. Tosupport operations with the spacecraft inclean area, the office premises areconnected with hall 119 by the requirednumber of cables run through hermeticallysealed inlets.

Work place for the SHM integration,processing and checks is also set up inarea B.

The AITB is equipped with the followingutilities: uninterrupted power supply (withthe same characteristics as both in Russiaand the US), heating, air conditioning andventilation, water supply, sewer, securityand fire alarms, and variouscommunication systems.

6.3 Spacecraft Fuelling Station

S fueling station is located in closevicinity to the AITB and is used for fillingspacecraft with liquid propellant andgases.

Fuelling station is equipped with followingsystems:

oxidizer filling system;

fuel filling system;

fuelling remote control system;

system for collection and incinerationof propellant vapors and sewage;

fuelling equipment neutralizationsystem;

temperature and humidity controlsystem;

vacuumization system;

gas control system;

fire fighting system;

set of scales.

Depending of the SC fuellingrequirements, the propellant componentsundergo necessary processing:

filtration through 20 or 5 micron filters;

providing required temperature andhumidity of propellant;

saturation with nitrogen or helium; and

degassing.

The fueling station is a heated 97-m longand 41-m wide building. The building isdivided into three sections: section No. 1 filling with fuel, section No. 2 - filling withcompressed gas, and section No. 3 filling with oxidizer.


The external doorways and openingsbetween the sections have hermetic seals rolling hermetic doors 5.5 meters longand 7 meters high. The sections areequipped with 15-ton capacity bridgecranes. For the purposes of transportation,there is a rail line running through thefilling rooms. The temperature in the fillingrooms is kept at +50C to +350C. Fillingrooms are equipped with clean tents,which provide the cleanliness class100,000 in accordance with US FederalStandard 209 A.

6.4 Launch Complex

Launch complex is a combination ofservice facilities, systems and lines, whichensure accomplishing of the followingtasks:

Achieving launch readiness status forthe Dnepr LV and SC;

Continuous and periodical automatedremote control of LV, SC and silolauncher equipment parameters; and

LV launch.

The launch complex includes:

silo launchers;

launch control center (LCC);

standard internal power supply system;

communication and control cable linesrunning between sites;

Figure 6.4-1 Dnepr Launch ComplexFacilities (Sites 106 and 109)

roads and technical facilities availablebetween sites.

If active electrical interface is used or theCustomer needs to monitor the SC statusup to launch, the underground Facility 25can be used for accommodation of the SCCOE. This facility is located in the closevicinity of the silo launcher and isconnected with it by a 80 m utility tunnel.Facility 25 is a reinforced concretestructure consisting of a number of rooms.In particular, a 67 m2 room (5.7 by 11.8 mand 2.4 m high) will be available for theCustomers COE.

Facility 25 has a 220/380V 3-phase 15kWt power supply system. Requiredtemperature (+ 15 - +250C) is maintainedby electric heaters.

Facility 25 can be equipped with variouscommunications means to communicatewith the LCC and other services both atBaikonur Cosmodrome and outside of it,as required by the Customer. ISDN line(64kbit/sec.) can be made available forFacility 25, if necessary.


In addition to that, the Customer may beoffered a so-called customer monitoringpost, located near the LCC at a distance ofabout 7 km from the launch site andFacility 25. These premises mayaccommodate customers personnel, whomonitor the SC parameters up to thelaunch. If the parameters being monitoreddiffer from the nominal values, theCustomer will have an opportunity tosuspend the pre-launch sequence no laterthan 3 minutes prior to launch in order tosave the spacecraft.

Figure 6.4-2 Transporter Erector

A transporter erector, equipped with a

system that allows to maintain certaintemperature and humidity conditions, isused for transportation of the SHMcontaining the SC from AITB to the silolauncher for integration with the launchvehicle.

The launch of the LV and control over thelaunch command execution are carried outvia wire communications by the remotecontrol system, the equipment of which islocated at the LCC.

If the launch is cancelled or postponed, orit is necessary to detach the SHM from thelaunch vehicle, the operations areconducted in a sequence reverse to theSHM/LV integration sequence.

In case of spacecraft on-board equipmentmalfunction, or if it is necessary to re-check the spacecraft at the SC processingfacility, operations of its delivery to theprocessing facility (Site 42 or 31) areperformed in a reverse sequence of SCintegration with the SHM.


7. Baikonur Operations Flow

Spacecraft and launch vehicle preparationoperations are divided into several stages,which may overlap depending on thespacecraft preparation requirements:

preparation of LV for integration withSHM;

SC processing;

preparation of SHM for mating with theSC;

SC / SHM integration;

transportation of SHM containing SCby transporter erector to the silolauncher;

SHM/LV integration and LV preparationfor launch;

pre-launch operations and LV launch.

Figure 7-1 shows work schedule forpreparation and launch of a SC byDnepr-1 LV (all dates are tentative).

Duration, days# Stage

1 Spacecraft processing

2 Preparation of SHM formating with the SC

3 SC/SHM integration andtransportation to launch silo

4 Preparation of LV forintegration with SHM

5 SHM/LV integration

6 Pre-launch operations andLV launch

Figure 7-1 Spacecraft Processing and Launch Schedule

20

9

7

22

1

6


7.1 Preparation of LV for Integration withSHM

LV preparation for integration with theSHM is done as follows:

installation of the launch vehicletransport and launch canistercontaining the 1st and 2nd stages intothe silo launcher;

equipping 3rd stage with telemetrysystem;

filling of 3rd stage with propellant;

mating 3rd stage and LV inside the silo;

set-up of work place for SC checks inFacility 25 (if active electrical interfaceis used);

checkout of the launch vehicle with theelectrical equivalent of the SHM, checkof SC/work place-Facility 25 line; and

fueling of stages 1 and 2 withpropellant components.

Figure 7.1-1 Loading of TLC ContainingLV 1st and 2nd Stages into Silo Launcher

Duration of LV preparation for integrationwith SHM is about 21 days. The fueled LVcan await integration with the SHM insidethe silo for an unlimited period of time(within its operational life).

7.2 SC Processing

Containers with the spacecraft and itsground processing equipment (GPE)arrive Baikonur by air or by rail, inaccordance with the launch campaignschedule agreed between the Customerand ISC Kosmotras, normally, no less than20 days prior to launch.

Container off-loading is done by liftinggear of the airplane, by road cranes orforklifts using pull ropes, cross-arms,pallets and other Customers devices.

Arrived containers with the spacecraft andGPE will be delivered to the SC/SHMprocessing facility for subsequentoperations by road or rail, on commoncarrier rail flatcars in shrouded condition toprotect from direct impact of precipitation.

Figure 7.2-1 Off-loading of SpacecraftContainer from the Plane


If necessary, the containers can betransported from Yubileiniy airport onspecial environmentally controlled railcars.At customers request, the groundtransport loads (g-loads, accelerations,temperature) can be controlled duringtransportation and the results madeavailable to the customer.

Figure 7.2-2 Spacecraft Processing

At SC and SHM processing facility, thecontainers with the spacecraft and GPEare off-loaded inside the AITB. GPE isdeployed at designated work places andthe spacecraft is brought into the cleanroom (clean room inside AITB at Site 42 orArea B inside AITB at Site 31).

Spacecraft processing is done based onthe spacecraft authority requirements,including its filling with propellantcomponents and gases at the fuellingstation, if necessary. To ensure therequired temperature environment duringthe spacecraft transportation from theAITB to the fuelling station, the spacecraftis delivered to the fuelling station inside itscontainer on a special environmentallycontrolled railcar.

The customer may use its own propellant(hydrazine).

7.3 Preparation of SHM for Mating withSC

Upon completion of the spacecraftprocessing (if processing is done at Site42), the SHM or EPM is delivered to theAITB by a special road vehicle.Upon delivery, the SHM is off-loaded andthe fairing is detached.

The SHM or EPM is brought inside theclean room by a special rail trolley.Adapter(s) is prepared for mating with thespacecraft. Spacecraft adapter(s) may bedetached from SHM or EPM and placedon special supports.

7.4 SC / SHM Integration

Counterparts of separation mechanismsare mounted on the spacecraft.


Spacecraft are mounted on adaptersmanually or by telphers available at cleanroom; spacecraft electrical checks andbattery charging (if necessary) areconducted. Then the spacecraft with theiradapters attached are mated with theSHM or EPM.

Final assembly of SHM or EPM is carriedout. SHM or EPM is then taken out of theclean room by a rail trolley. Fairing isattached to the SHM.

SHM or EPM is loaded into a special roadvehicle and transported to the AITB at Site31 for SHM final assembly, its electricalchecks and preparation for integration withthe launch vehicle.

If the spacecraft is processed at Site 31,

Figure 7.4-1 SHM Integration

the assembly of the SHM (with or withoutEPM) is conducted at the Area B of theAITB at Site 31 following the samesequence as for the Site 42, except for thetransportation of the SHM (EPM) betweensites.

7.5 Transportation of SHM ContainingSC by Transporter Erector to the SiloLauncher

Assembled and processed SHM is loadedinto the transporter-erector and deliveredto the launch complex. During thetransportation, the required temperatureand humidity conditions are maintainedaround the spacecraft by means oftransporter erector climate control system.SHM delivery time to the launch complexdoes not exceed 3 hours.

Figure 7.5-1 SHM Loading intoTransporter Erector


7.6 SHM/LV Integration and LVPreparation for Launch

Upon arrival at the launch complex, theSHM, by means of transporter erector, ismated with the LV installed inside the silo.SHM/LV mating time does not exceed 3hours.

Electrical checks of the LV with SHM, dataanalysis, final launch preparationoperations are performed. Time periodfrom the completion of SHM/LV mating tolaunch does not exceed 6 days.

If necessary, during LV preparation forlaunch, continuous (periodical) monitoringof the spacecraft from Facility 25 can beorganized. Monitoring stops 30-40 minutesprior to launch since all personnel must beevacuated from danger zone to the vicinityof the LCC. 24-hour stand-by duty can beorganized at the launch complex tosupport operations with the spacecraftCOE.

Figure 7.6-1 SHM/LV Integration InsideSilo Launcher


7.7 Pre-launch Operations and LVLaunch

The LV launch is conducted from the LCC(Site 111/2). The launch command isissued at the pre-determined moment oftime, and the following operations areperformed prior to the launch commandissuance:

final operations are conducted at theLCC (3 hours);

2 hours prior to launch, the preparationof the telemetry system groundequipment begins;

Transporter Erector

Space Head Module

Silo Launcher

Figure 7.6-2 SHM Loading into Silo Launcher


1.5 hours prior to launch the silo dooris opened and mobile equipment in thevicinity of silo is evacuated;

1 hour prior to launch, the on-boardtelemetry system is activated (for about10 minutes) and the telemetry datareceipt by the ground stations isverified;

20 minutes prior to launch, the on-board telemetry system is activatedagain and the personnel areevacuated;

3 minutes before launch by acommand from the launch control post,the on-board telemetry system startsgetting power from the LV on-boardpower supply system, the groundstations start receiving, recording andprocessing the actual pre-launch andflight telemetry;

LV is launched.

After the launch, the launch crew andtelemetry system crew return to the silolauncher and conduct final operations withthe telemetry system checkout equipmentand examination of the silo launcher. Figure 7.7-1 Dnepr LV Liftoff


8. SC/LV Interfaces

8.1 Mechanical Interface

Spacecraft is attached to the launchvehicle adapter by means of pyro-devices.The number and type of pyro-devices iscontingent on a number of parameters(spacecraft weight, the diameter on whichthe attachment points are located, etc.)and is agreed upon with the customer forthe specific spacecraft.

For some types of spacecraft, theadapters that were used in previouslaunches, can be applied. They includeadapters for 10-20 kg spacecraft, 50-70 kgspacecraft and 300-400 kg spacecraft.These adapters can serve as prototypesfor the development of new ones.

Typical design of the spacecraft/standardadapter attachment point is shown inFigure 8.1-1.

To verify the spacecraft separation fromthe Space Head Module, separationswitches mounted on the adapter areused.

To disconnect electrical connectors thatensure the LV/SC electrical links,separation mechanisms are used, whichare activated prior to the operation of theseparation system.

Pyro-devices, separation switches andmechanisms used, have undergone allnecessary ground and flight testing andare highly reliable.

Separation system activation equipmentand cables are installed on the SpaceHead Module and adapter in compliancewith all operational requirements, whichprecludes any damage to or collision withthe elements being separated and

satisfies high reliability requirements ofelectrical command relay that wasconfirmed by ground and flight testing.

If necessary, and subject to mutualagreement, the adapter can be supplied tothe spacecraft authority by ISC Kosmotrasand SDB Yuzhnoye for fit-check testing.

It should be noted that the separationsystems used for SC/Dnepr-1 LVseparation have no spring pushers, sincethe spacecraft separation is done bytaking away the upper stage from thespacecraft by means of the upper stagemotor throttled-back operation, whichensures minimum disturbances on thespacecraft during separation.

Angular stabilization errors of the launchvehicle 3rd stage with the Space HeadModule at the moment of issuing thespacecraft separation command, do notexceed:

+ 1.5 degrees for angles of pitch, yawand roll; and

+ 0.5 degrees per second for rates ofpitch, yaw and roll.

Spacecraft disturbances due to process ofseparation are dependant on the inertialcharacteristics of the spacecraft and the3rd stage with the Space Head Module(including the spacecraft), on the type,number, location and characteristics of thespacecraft/space head module attachmentjoints. Spacecraft (with the weightexceeding 300 kg) angular rates afterseparation due to stabilization errors anddisturbances induced by the separationprocess are as follows:

x 2.0 degrees per second;y 3.0 degrees per second;z 3.0 degrees per second.


Analysis of the relative motion of thespacecraft being separated and SHMelements confirms the spacecraft

separation safety, i.e. indicates no collisionamong the spacecraft themselves and withthe SHM elements.

* - 20 by 20 mm seats, where separation switches will rest, should be available on the spacecraft.

Figure 8.1-1 Typical Design of Spacecraft/Standard Adapter Attachment Point


8.2 Electrical Interface

Two types of SC/LV electrical interface arepossible when launching a spacecraft onDnepr-1 LV:

Passive electrical interface noelectrical connections between LV andSC. No power is supplied to thespacecraft during ascent. Uponspacecraft separation, its separationswitch is activated and its power supplysystem turns on. The LV hasspacecraft separation sensors which,upon separation, generate a telemetrysignal that is transmitted to the ground;and

Active electrical interface electricalconnections available between SC andLV that are used prior to (SC-SHM-LV-TLC-SC COE) and during launch (SC-SHM-LV) as well as during injectioninto orbit (SC-SHM).

The type of the electrical interface isagreed upon between the Customer andISC Kosmotras. The above interface typesdiffer by their complexity and cost of theirtechnical realization.

SC/LV electrical links ensure the following:

transmission, if necessary, of telemetrydata on the status of the SC on-boardsystems during ascent as well asduring electrical checks of the SCintegrated with the LV inside the silolauncher;

issuance of commands by the LVcontrol system to the SC on-boardinstrumentation during ascent as wellas during electrical checks of the SCintegrated with the LV inside the silolauncher;

supply of power for SC batterycharging, maintaining requiredtemperature and humidity conditions ofthe SC bus and for the SC commandand control interface, prior to launchwhen the SC is integrated with the LVinside the silo launcher.

Existing through circuits of the LV controlsystem and TLC that can maintain about20 various transmission lines are used forLV/SC and LV/SC checkout equipmentelectrical links. Diagram ofSpacecraft/Spacecraft ground equipmentlinks is shown in Figure 8.2-1.

Through circuits run via umbilicalconnectors on the LV frame, which areruptured prior to launch 3.5 secondsbefore LV starts moving inside the TLC.Spacecraft ground equipment isconnected to the LV and TLC throughcircuits by means of additional connectorsmounted on the TLC and cables runningfrom the silo to Facility 25.

A 32-pin electrical connector (8) isavailable at LV/SHM interface, the pins ofwhich have 0.75 mm2 wires running from a3-pin electrical connector (8) that isinstalled at ground/LV side connectorboard and allows to pass the totalamperage of up to 50 A. On the SHM,between the 8 connector and twospacecraft umbilical connectors, a newcable will be run for battery charging andmaintaining the spacecraft on-boardequipment required temperature andhumidity conditions.

In addition to that, LV/SHM interface hasthree more 102-pin electrical connectors(1, 2, 3), some pins of whichhave 0.2 mm2 and 0.35 mm2 wires forthrough circuits running from electrical


connectors (2, 3, 4) installed atthe side connector board. These circuits(about 65 of them) can be used forspacecraft/ground equipment link. For thispurpose, a new cable will be run from thespecified connectors to the third umbilicalconnector of the spacecraft.

If the characteristics (wire cross-section,shielding, etc.) of the above-mentionedstandard LV electrical links are notsufficient to ensure the required electricalinterface with the SC, an additional cablemay be inserted into the electricalconnectors (2, 3, 4) of the sideconnector board. At the LV/SHM joint, theoutput connectors of this additional cablewill be mated with the SHM cables thatservice the Spacecraft.

It is desirable to install on the spacecraftseveral hermetically sealed subminiatureplug connectors used on SS-18 (standarddesignation 50 .364.046), the pins of which are soldered with0.12 0.35 mm2 wires. The plug weight is38 grams (with the case) and 20 grams(without the case). Maximum amperage is75 A. This connector can withstand lowpressure (down to 10-6 mm of Mercury)and pressure difference of up to 1atmosphere. This connector has astandard SS-18 no-impact disconnectiondevice.

All electrical connectors used in throughcircuits are standard proven devicessuccessfully tested by a great number ofSS-18 launches. Other types of umbilical

Figure 8.2-1 Schematics of SC/SC COE Transit Electrical Links


connectors can be installed on theSpacecraft. Their counterparts are to beprovided to ISC Kosmotras for fabricationof the appropriate SHM cables.

When the SC is injected into the requiredorbit, the LV control system, if necessary,issues a reference command to the SCinstrumentation. Upon 0.1 secondfollowing the issuance of the command,the electrical connectors with SC aredisconnected and 0.5 second after theissuance of the command, the SC/LVseparation occurs.

If active electrical interface is used or it isnecessary to monitor the SC status up tothe moment of launch, the SC COE can beinstalled in underground Facility 25 locatedin close vicinity of the silo launcher.

During electrical checks of the SCintegrated with LV inside the silo, thefollowing parameters of SC/LV interfaceare monitored:

connection of cables running from LVto SC and ground equipment;

issuance of signal INITIAL STATUSOF LV CONTROL SYSTEM;

resistance of power lines supplyingsignals from the LV control system tothe SC instrumentation; and

issuance of reference signals from theLV control system to the SCinstrumentation.

The list of commands (signals) required formaintaining interface with the LV controlsystem during LV preparation for launch,launch and flight, as well as sequence andparameters of such interface, types ofconnectors and number andcharacteristics of electrical links areagreed upon between ISC Kosmotras andthe spacecraft authority.


9. Spacecraft Environments

9.1 Stiffness Criteria (FrequencyRequirements)

The spacecraft should be designed with astructural stiffness, which ensures that thevalues of fundamental frequency of thespacecraft, hard mounted at theseparation plane, are not less than:

20 Hz in the longitudinal axis; and

10 Hz in the lateral axis.

If it is not possible to comply with theabove requirements, SDB Yuzhnoye willcarry out an additional analysis of the LVdynamic characteristics and loads that willtake into account the spacecraftfundamental frequencies.

9.2 Quasi-static and Dynamic Loads

Tables 9.2-1 and 9.2-2 contain quasi-staticand dynamic components of accelerationsthat act on the SC/LV interface during theground handling, launch and in-flight.

Spacecraft dimensioning and testing musttake into account safety factors, which aredefined by the spacecraft authority, butshould be no less than the values givenbelow:

2.0 for ground handling;

1.5 during launch while LV is movinginside the TLC;

1.3 during launch after the LV exitsfrom the TLC;

1.3 during the LV flight.

The spacecraft should remain operableafter the effect of the above accelerations.

9.3 Vibration Loads

Described below are vibrations acting onthe Spacecraft attachment points duringthe LV flight. Two types of vibrations areas follows:

Harmonic oscillations; and

Random vibrations.

The harmonic oscillations arecharacterized by the amplitude of vibro-accelerations and frequency. Theparameters of harmonic oscillations aregiven in Tables 9.3-1 and 9.3-2.

The random vibrations are characterizedby spectral density of vibro-accelerationsand the duration of influence. The randomvibration parameters are given in Table9.3-3.

The random vibrations are spatial withapproximately equal intensity of vibro-accelerations in each of the threerandomly selected mutually perpendiculardirections.

The values of amplitude and spectraldensities are given in the extreme octavepoints. The change of these values withinthe limits of each octave is linear in thelogarithm frequency scale.


Table 9.2-1 Accelerations at SC/LV Interface during Transportation

AccelerationLoad Source

Longitudinal (X) Lateral (y) Lateral (z)

SHM Transportation 0.4 -1.00.7 0.5

Table 9.2-2 Maximum Quasi-static and Dynamic Accelerations at SC/LV Interface

AccelerationLoad Source Longitudinal (X) Lateral (y, z)

LV movement inside TLCAfter LV exit from TLC

2.50.71.0

0.30.8

1st stage burn:Maximum dynamic headMaximum longitudinal acceleration

3.00.57.50.5

0.50.50.10.5

2nd stage burn maximumlongitudinal acceleration 7.80.5 0.2

3rd stage burn -0.3...-0.5 0.25

Notes to Tables 9.2-1 and 9.2-2:

Lateral accelerations may act in any direction, simultaneously with longitudinalones;

The above values are inclusive of gravity force component;

Dynamic accelerations are preceded by "" symbol;

The above values are correct for the spacecraft complying with the fundamentalfrequency requirements contained in paragraph 9.1.


Table 9.3-1 Amplitude of Harmonic Oscillations at SC/LV Interface. LongitudinalAxis (X)

Frequency sub-band, Hz 5-10 10-15 15-20

Amplitude, g 0.5 0.6 0.5

Duration, sec. 10 30 60

Table 9.3-2 Amplitude of Harmonic Oscillations at SC/LV Interface. Lateral Axes (Y, Z)

Frequency sub-band, Hz 2-5 5-10 10-15

Amplitude, g 0.2-0.5 0.5 0.5-1.0

Duration, sec. 100 100 100

Table 9.3-3 Spectral Density of Vibro-accelerations at SC/LV Interface

Load Source

Frequency sub-band, Hz Liftoff, LV flightsegment where M=1,qmax

1st stage burn (except for LVflight segment where =1,qmax), 2nd stage burn, 3rdstage burn

Spectral Density, g2/Hz

20-40 0.007 0.007

40-80 0.007 0.007

80-160 0.007-0.022 0.007

160-320 0.022-0.035 0.007-0.009

320-640 0.035 0.009

640-1280 0.035-0.017 0.009-0.0045

1280-2000 0.017-0.005 0.0045

Root Mean Square Value, , g 6.5 3.6Duration, sec. 35 831


9.4 Shock Loads

Shock loads are wide-band, fadingprocesses and are characterized by theshock spectrum and the duration of action.

The activation of the separation pyro-devices is a source of the vibro-pulseloads at the spacecraft attachment points(the duration of shock process is up to0.1 sec). The shock spectrum values aregiven in Table 9.4-1. They are accurate forthe Q=10 and for each of the threerandomly selected mutually perpendiculardirections. The change of the shockspectrum values versus frequency withineach sub-band is linear (in the logarithmfrequency scale and shock spectrumvalues).

9.5 Acoustic Loads

The sources of acoustic loads are:

1st stage motor burn;

frame surface pressure fluctuations inthe turbulent boundary layer.

The acoustic loads are characterized bythe duration of action, integral level of thesound pressure within the frequency bandof 20-8,000 Hz, and the levels of soundpressure within the octave frequencyband with the mean geometricfrequencies of 31.5; 63; 125;; 2,000;4,000; 8,000 Hz.

Table 9.4-1 Shock Spectrum at Spacecraft Attachment Points

Frequency subband, Hz

30-50

50-100

100-200

200-500

500-1000

1000-2000

2000-5000

Load Source

Shock Spectrum Values, g

Numberof shockimpacts

Separation of fairing, 3rdstage and neighboringspacecraft

5-10

10-25

25-100

100-350

350-1000 1000 1000 *

Separation of SC 5-1010-25

25-100

100-350

350-1000 1000

1000-3000 1

Note: * - number of shock impacts is contingent on a number of spacecraft installed inthe SHM.


9.6 Temperature and HumidityConditions and Thermal Effect onSpacecraft

During operations with the spacecraft atSC processing facility, the air temperaturearound the spacecraft is maintained within21 - 270C, with relative humidity of notmore than 60%.

During SC/SHM integration at AITB, theair temperature is maintained within 5 -350C, with relative humidity of not morethan 80%.

When transporting the SHM to SHMprocessing facility and to the launch silo,the temperature inside the Transporter-Erector is within 10-250C with relativehumidity of no more than 80%.

During operations with the SHM at SHMprocessing facility, the air temperaturearound the spacecraft is maintained within5 - 350C, with relative humidity of not morethan 80%.

When loading the Space Head Moduleinto the launch silo and mating it with theLV, the SHM is affected by thetemperature within 0-450C during the timeperiod of no more than 30 minutes andwith the temperature within 5 - 350C duringthe time period of no more than 5.5 hours,with the relative humidity being no morethan 80%.

When the SHM is inside the silo, thetemperature inside the silo is within therange of 5 - 250C with the possible short-term increase of up to 350C and relativehumidity is of no more than 80%, and thetemperature around the spacecraft is

Table 9.5-1 Acoustic Loads

Mean Geometric Frequency of Octave Frequencyband, Hz

Level of Sound Pressure, dB

31.5631252505001000200040008000

125132135134132129126121115

Integral Level of Sound Pressure, dB 140

Duration, sec. 35


within the range of 5 - 300C with therelative humidity being no more than 70%.

Spacecraft heat emission while on the LVinside the silo and in-flight were not takeninto account.

Thermal flux acting on the spacecraft fromthe inner surface of gas-dynamic shieldwill not exceed 1,000 Wt/m2.

9.7 Pressure Underneath LV Fairing

Pressure change inside the fairingenvelope during the ascent phase is givenin Figure 9.7-1.

The maximum rate of in-flight pressurechange inside the fairing envelope doesnot exceed 0.035 kgf/(cm2 per sec.),except for transonic phase of flight where

a short term (2-3 seconds) increase up to0.035 kgf/(cm2 per sec.) is possible.

Data contained in this section may bespecified for each specific mission.

9.8 Gas-dynamic Effect on Spacecraft

Following separation from the Space HeadModule the spacecraft encounters a shortterm impact (several seconds) of the 3rdstage motor plume.

All combustion products (composed of: N2 28%, H2 27%, H2O 21%, CO2 18%,CO 6%) are in gaseous state; solid orliquid phases are not present.

Parameters of the 3rd stage motor plumeaffecting the spacecraft are given in Table9.8-1.

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

0 20 40 60 80 100 t,sec.

P, kgf/cm2

Figure 9.7-1 Pressure Change Rate inside Payload Envelope


Spacecraft surface contamination due tosedimentation of solid or liquid particlesdoes not occur, since they are not presentamong the 3rd stage motor combustionproducts.

Spacecraft surface contamination due toH2O vapor condensation does not occur,since the maximum gas pressure on thespacecraft surface (stagnation pressure) issignificantly (dozens of times) lower thanH2O saturated vapor pressure at thespacecraft surface temperature, which, atthis moment, normally exceeds 2730K.

3rd stage motor plume impact on thespacecraft is insignificant. Integral thermalflux on the spacecraft surface duringseparation will not exceed 5 Wt perhour/m2. Actual heating of 1 mm thicknessspacecraft shell made of AMg-6 alloy is 2-40.

3rd stage motor plume may cause lightdisturbances of the spacecraft motion.Torque may be induced that is dependanton the spacecraft position with respect tothe X axis of the launch vehicle andspacecraft inertia characteristics.

Data on the spacecraft launched byDnepr-1 LV and history of theirsubsequent operation confirm the absenceof 3rd stage motor plume impact onspacecraft operability, including thespacecraft equipped with high resolutionoptics.

9.9 SC/LV Electromagnetic Compatibility

The engineering solution to use EPM orgas-dynamic shield provides for about 30dB noise shielding within the frequencyrange of 10 kHz 30 GHz. Apart formthat, up to about 275th second of flight, theLV fairing provides for additional 10 20dB noise shielding within the frequencyrange of 10 kHz 1000 MHz.

Maximum residual levels of LV on-boardsystem (radioelectronics) RF emissionsthat penetrate the EMP (GDS) and haveeffect on the spacecraft during the activephase of LV flight prior to and after theEMP cover (GDS) separation, are given inTable 9.9-1.

Table 9.8-1 Maximum Values of 3rd Stage Motor Plume Parameters Affecting Spacecraft

, m 0 1.8 3.3 4.5 6 7.5 9 15 20 30 50

t,sec. 0 1.2 1.6 1.9 2.2 2.5 2.7 3.5 4.0 4.9 6.3

0, kgf/m2 0.2 6.27 4.7 7.3 11 8.2 6.6 3.2 1.84 0.84 0.3

, kgf/m2 0.05 0.048 0.034 0.025 0.017 0.012 0.0095 0.0038 0.002 0.00077 0.00023

, 254 252 238 226 213 200 193 165 149 127 104

V, m/sec. 3670 3671 3680 3687 3696 3704 3709 3726 3736 3750 3764

Where: x - distance (along X axis) from EPM attachment plane; P0 - gas stagnation pressure; P, T, V - pressure, temperature and velocity respectively of undisturbed gas flow.


Maximum levels of industrial noiseinduced by LV control system equipmentthat affect the spacecraft during activephase of LV flight within the frequencyrange of 10 kHz 1 GHz, will not exceed5.6 mV prior to and 35 mV after the EMPcover (GDS) separation.

Maximum levels of electromagneticemissions (radio and industrial noise) forLV on-board systems are given in Table9.9-2.

Maximum levels of electromagneticemissions generated by spacecraft and

Table 9.9-1 Maximum Residual Levels of RF Emissions

Electormagnetic Field Strength, V/mFrequency Band

Prior to EPM (GDS) CoverDrop

After EPM (GDS) CoverDrop

10 kHz 140.4 MHz140.4 144.4 MHz144.4 1000.5 MHz1000.5 1004.5 MHz1004.5 2500 MHz2500 2855 MHz2855 2865 MHz2865 30000 MHz

1.410-2

0.431.410-2

0.381.410-2

5.410-2

1.25.410-2

2.210-2

6.92.210-2

62.210-2

0.12190.12

Table 9.9-2 Maximum Levels of Electromagnetic Emissions

Frequency Band Electormagnetic Field Strength10 kHz 125 MHz125 250 MHz250 1000 MHz1000 1050 MHz1050 1570 MHz1570 1620 MHz1620 2750 MHz2750 2900 MHz2900 7500 MHz7500 7600 MHz7600 30000 MHz

70 mV/m10 V/m70 mV/m10 V/m70 mV/m10 V /m70 mV/m50 V/m70 mV/m10 V/m70 mV/m


spacecraft electric checkout equipment(ECOE) at the SC/LV and LV/spacecraftECOE interfaces (at a distance of 1 mfrom spacecraft and its ECOE), from thebeginning of SC/LV integration and untilspacecraft separation from LV plus 1minute, must be 10 dB less than thevalues given in Table 9.9-2.

During preparation and launch ofspacecraft that has active electricalinterface with LV, the spacecraftinstrumentation is also affected, along theelectrical circuits of SC/LV interface, byelectromagnetic interference with thelevels of up to 100 dB/V within thefrequency range of 30 Hz 100 MHz forfeed and control circuits, and up to 60dB/V within the frequency range of 30 Hz 30 GHz for data circuits.

Instrumentation of SC, in case of activeelectrical interface, should not generateelectromagnetic emissions in the electriccircuits of the SC/LV interface over 60dB/V within the frequency range of 30 Hz 100 MHz for feed and control circuits,and over 40 dB/V within the frequencyrange of 30 Hz 30 GHz for datacircuits.

Density of electromagnetic fieldsgenerated by Cosmodrome electronicequipment at processing facilities androutes of transportation of SC, SHM andLV do not exceed 10 V/m for frequencyrange 10 kHz 30 GHz, except for thefrequency range of 1570 - 1620 MHz,where maximum allowable level ofexternal electromagnetic interference is 30mV/m (for active or shut radio receivers ofspacecraft). This is similar to bothEuropean and US standards.

Coordination of efforts to ensureelectromagnetic compatibility of SC, LVand range systems at all stages of SC/LVintegration is done under a special EMCprocedure for a specific mission that isdeveloped at the preliminary integrationphase.

9.10 Spacecraft Tests Required to MeetDnepr LV Launch Services Requirements

The customer shall demonstrate that thespacecraft meets the requirementsdetailed in the entire section 9 of thisUsers Guide, by means of analyses andground tests.

For spacecraft qualification andacceptance, sinusoidal, shock and randomtests are mandatory.

A test plan established by the spacecraftauthority describing the tests, which areexecuted on the spacecraft, shall beprovided to SDB Yuzhnoye.

After completion of the tests, the testresults report shall be submitted to SDBYuzhnoye.

10. Ground Qualification Tests

To verify the ability to integrate thespacecraft with the Dnepr launch vehicleand to confirm the operability ofspacecraft/launch vehicle mechanical andelectrical interfaces, a ground qualificationtest program may be provided thatincludes spacecraft fit-check testing,vibration testing of spacecraft and SHMstructural elements and spacecraftseparation system tests.

The objective of qualification testing is toconfirm the following:


operations with spacecraft provided forby spacecraft authority are easy tohandle by personnel and ensure no-collision integration with launch vehicle;

engineering solutions to protect thespacecraft from damage are sufficient;

vibration strength at spacecraftattachment points is sufficient towithstand loads on the spacecraftduring LV launch and flight;

separation system remains operablefollowing the impact of vibration loads;

complete separation alongspacecraft/adapter joints is achieved;

motion parameters of the spacecraft

being separated meet therequirements and shock loads that actupon activation of spacecraftseparation pyros are within theallowable limits.

Qualification tests are performed at SDBYuzhnoye facilities with participation ofspacecraft authority specialists. The scopeof qualification testing is to be agreedupon with the spacecraft authority.

Necessary process related equipment andspacecraft dummies with actualmechanical and electrical interfaces to befurnished by spacecraft authority are usedfor qualification testing. Based on theresults of qualification testing, anacceptance certificate authorizing thelaunch of a specific spacecraft on DneprLV is released.


11. Telemetry and Tracking

Telemetry system is based on LV on-board radio telemetric system with datacapacity of 512 Kbit/sec., which allows toregister both slowly and rapidly changingparameters.

During spacecraft injection, a number ofphysical process parameters areregistered, which determine the properfunctioning of the LV systems and units,characterize the LV/SC separationprocess as well as measure dynamic andheat impact on the spacecraft during itsorbital injection. For this purpose, the LVon-board telemetry system is equippedwith initial signal transformers andnormalizers.

Data capacity of the LV telemetry systemallows, if necessary, for recordinginformation supplied by the spacecraft on-board instrumentation within the broadsampling frequency range of 25 Hz 8kHz and output voltage of 0-6 V.

Specific requirements for telemetryreading and interface will be definedduring SC/LV integration concept work.

Pre-launch checks of the LV on-boardtelemetry system are carried out by meansof available standard ground checkoutequipment.

Motion parameters of the LV center ofgravity are determined by GPS system,which transmits obtained data throughtelemetry channels.

When launching spacecraft on Dnepr LVfrom Baikonur eastward (orbit inclinations50.5 and 64.5 degrees), the telemetry datatransmitted from LV through two channels(to ensure appropriate transmissionreliability) is registered by Baikonur groundtracking stations and ground trackingstations of Russian Federation locatedalong the LV flight trajectory.

When launching spacecraft from Baikonursouthward (orbit inclinations 87.3 and 98.0degrees), the telemetry data issubsequently registered by a mobileground tracking station located in Omannear the town of Salalah. This groundstation relays to Baikonur data collectionand processing center a certain amount oftelemetry data in real-time scale, includingthe parameters of spacecraft initial motion.

Figure 11-1 Tracking Station


12. Analysis of Flight Results

Analyses of all LV system functioning, ofthe loads the spacecraft encounteredduring injection are carried out based onthe results of the telemetry measurementstaken during the LV pre-launchpreparation and flight.

The scope of analyses required isdetermined jointly with the customer in theprocess of approval of the statement ofwork for the spacecraft launch.

The analysis of the flight results is carriedout in the following way:

real-time telemetry data on a limitednumber of LV system statusparameters (command issuance andexecution, propulsion unit operationand LV flight stabilization);

determination of spacecraft separationevent (10 minutes after the separation);

determination of spacecraft actualseparation time (1.5 hours after thespacecraft separation event);

preliminary data on spacecraft initialmotion parameters (2 hours afterspacecraft separation);

full set of data on spacecraft orbitalparameters (1 day after spacecraftseparation);

comprehensive analysis of all the datareceived in order to evaluate the LVsystem and sub-system functioning, toget their quantitative values, to assessspacecraft injection environments andto determine the probable reasons ofmalfunctions if there were any (1month after launch).

Figure 12-1 Flight Data ProcessingCenter


13. Range Safety

ISC Kosmotras possesses certainexpertise and proven safety procedures,both organizational and technical, forlaunch campaigns. All works andoperations are conducted in strictaccordance with approved plans anddocuments.

Our company guarantees that theCustomers representatives have fullcontrol over the access to the spacecraft,as far as it is technically feasible. Jointoperations of Customers and Providersspecialists are conducted in compliancewith the approved procedures. Followingthe completion of operations in clean roomfacility (SC/adapter integration, assemblyof EPM), the EPM cover attachment pointsare sealed by Customers representatives,this fact is registered in a specialcertificate and recorded by video andphotography. Following this procedure, thedirect access to the spacecraft becomesimpossible. If necessary, visual andremote monitorin

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