REQUEST FOR PROPOSAL · 3.1 isro ka hts-1 For HTS-1 the contractor is fully responsible for all...

ISRO/ISAC/HTS/RFP/001/06VER-01

REQUEST FOR PROPOSAL

ISRO Ka BAND HIGH THROUGHPUTSATELLITES

December 2015

ISRO Satellite CentreBangalore

December 2015ISRO CONFIDENTIAL PROPRIETARY

ISRO/ISAC/HTS/RFP/001/06VER-01 CONTENTS

CONTENTS

Schedule A Statement of Work A1-A25

Schedule B Technical Specifications Requirement B1-B89

Schedule C Assembly, Integration, Testing & Satellite Storage Plan C1-C7

Schedule D Product Assurance Requirements D1-D6

Schedule E Launch Interfaces & Launch Support Services E1-E5

Schedule F Mission & Post launch Support Services F1-F7

Schedule G Commercial & Contractual Requirements G1-G24

Annexure - A Compliance Matrix CM1-CM19

Annexure - B Abbreviations Ab1-Ab5

ISRO

REQUEST FOR PROPOSAL (RFP)FOR

ISRO Ka BAND HTS

SCHEDULE A

STATEMENT OF WORK

ISRO/ISAC/HTS/RFP/001/06VER-01 SCHEDULE - A

STATEMENT OF WORK

December 2015 Page A-1ISRO CONFIDENTIAL PROPRIETARY

TABLE OF CONTENTS

SECTION I ..................................................................................................................................31 INTRODUCTION .............................................................................................................52 SCOPE ............................................................................................................................5SECTION II .................................................................................................................................71 INTRODUCTION .............................................................................................................92 FLIGHT SATELLITES .....................................................................................................92.1 DEFINITIONS .................................................................................................................102.1.1 SPACECRAFT ................................................................................................................102.1.2 COMPATIBLE LAUNCH VEHICLE ..................................................................................102.1.3 OPERATIONAL LIFE ......................................................................................................102.1.4 DESIGN LIFE .................................................................................................................102.1.5 PROPELLANT LIFE ........................................................................................................102.1.6 BEGINNING OF LIFE (BOL) ............................................................................................102.1.7 END OF LIFE (EOL) ........................................................................................................103 RESPONSIBILITY .........................................................................................................113.1 ISRO KA HTS-1...............................................................................................................113.2 ISRO KA HTS-2...............................................................................................................113.3 ISRO KA HTS-1 AND ISRO KA HTS-2 ...............................................................................113.4 RISK MANAGEMENT .....................................................................................................123.5 ANOMALIES HANDLING AND DOCUMENTATION ..........................................................123.6 EXPORT AUTHORIZATIONS...........................................................................................124 REALIZATION PLAN & CUSTOMER’S TECHNICAL VISIBILITY AND ACCESS ...........124.1 REALIZATION PLAN ......................................................................................................124.2 TIMELINE AND PROCUREMENT ....................................................................................134.3 ISRO PARTICIPATION ....................................................................................................134.4 HERITAGE .....................................................................................................................134.5 ACCESS TO WORKPLACE ..............................................................................................144.6 SPACECRAFT LIFE AND MASS .......................................................................................144.6.1 SPACECRAFT LIFE.........................................................................................................144.6.2 SPACECRAFT MASS ......................................................................................................145 HARDWARE AND SOFTWARE FOR GROUND SYSTEM VALIDATION TESTING

AND SATELLITE OPERATIONS ....................................................................................145.1 TCR SUITCASE ..............................................................................................................155.2 TELEMETRY DATA........................................................................................................155.3 SATELLITE SIMULATOR ................................................................................................155.4 PAYLOAD RECONFIGURATION SOFTWARE TOOL .........................................................165.5 STEERABLE ANTENNA CONTROL SOFTWARE TOOL .....................................................165.6 ADCS RE-PROGRAMMING TOOL.................................................................................... 165.7 TRAINING .....................................................................................................................165.8 WARRANTY AND MAINTENANCE POST ACCEPTANCE ..................................................16


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5.9 TELECOMMAND ENCRYPTION......................................................................................176 LAUNCH VEHICLE COMPATIBILITY ..........................................................................177 GROUND CONTROL COMPATIBILITY ........................................................................178 PRODUCT ASSURANCE REQUIREMENTS.................................................................... 179 SATELLITE REVIEWS AND SCHEDULE .......................................................................189.1 EQUIPMENT QUALIFICATION STATUS REVIEW (EQSR) .................................................189.2 DESIGN REVIEWS ..........................................................................................................189.2.1 PRELIMINARY DESIGN REVIEWS (PDR)......................................................................... 189.2.2 CRITICAL DESIGN REVIEWS (CDR)................................................................................ 189.3 MANUFACTURING, ASSEMBLY, INTEGRATION AND TEST REVIEWS .............................189.3.1 SATELLITE QUALIFICATION STATUS REVIEW (SQSR) / ENVIRONMENTAL TEST

READINESS REVIEW (ETRR) ..........................................................................................199.3.2 FLIGHT MODEL COMPLETION REVIEW (FMCR)............................................................. 199.3.3 TEST REVIEWS ..............................................................................................................199.3.4 IN ORBIT ACCEPTANCE REVIEW (IOAR)........................................................................ 1910 NON DELIVERABLE EQUIPMENT................................................................................ 2011 SPARE PROVISIONS .....................................................................................................2012 SATELLITE STORAGE AND SAFE KEEPING ................................................................2013 SERVICES .....................................................................................................................2013.1 LAUNCH SUPPORT SERVICES ........................................................................................2013.2 MISSION SUPPORT SERVICES ........................................................................................21SECTION III ..............................................................................................................................231 PROGRAM MANAGEMENT ..........................................................................................252 MAJOR REALIZATION MILESTONES ..........................................................................25


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SECTION I

INTRODUCTION


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1 INTRODUCTIONISRO is playing a lead role in establishing a strong national communication infrastructure forsocietal, educational, industrial and entertainment sector applications. In order to enhance andupgrade this infrastructure, a new generation of spacecraft bus and payloads are required tobe inducted. In this direction, the design, development and realization of next generationstate of the art Ka-band high throughput satellites into the geostationary orbit will help infulfilling this requirement.

2 SCOPEThis Request for Proposal (RFP) defines the work to be performed by the Contractor for theSpacecraft program and defines all required deliverable hardware, software, documentationand services. All work shall be performed and completed in accordance with the terms andprovisions of the Agreement and its Schedules.

The following documents may be referred to in the text of this RFP, either by title orSchedule letter; either designates the applicable contractual document of the Contract:

Title Schedule

Statement of Work Schedule A

Technical Specifications Requirement Schedule BAssembly, Integration, Testing & Satellite Storage Plan Schedule C

Product Assurance Requirements Schedule D

Launch Interfaces & Launch Support Services Schedule EMission & Post launch Support Services Schedule FCommercial & Contractual Schedule G

A strong heritage of making high power, multi-beam Ka-band high throughput spacecraft isthe essence of this RFP. Any contractor proposing the spacecraft shall have a heritage of suchspacecrafts in orbit either on their own or in participation with other contractors.Documentary evidences to be provided to support heritage.

A detailed compliance matrix for the RFP shall be submitted by the contractor.

Contractor may note that, all information given by the contractors to customer will be keptconfidential.

In this document ‘Customer’ means ISRO and ‘Contractor’ means party quoting against RFP.

All program, technical documentation and meetings related to this program are required to bewritten and held in English.


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SECTION II

DELIVERABLES


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1 INTRODUCTIONThis chapter describes items to be delivered to customer, as well as the programme tasks tobe performed by the Contractor if selected under the Contract. The detailed requirements forthe tasks are defined below.

Lack of mention of any tasks which are required to perform this Contract will not be anexcuse for the non-completion of the above referenced task.

2 FLIGHT SATELLITESThe procurement is for two Ka Band High Throughput Satellites, named ISRO Ka BandHigh Throughput Satellite - 1 and ISRO Ka Band High Throughput Satellite - 2 foruse in a Geostationary Satellite Orbit (GSO). The overall functional requirements andtechnical specifications of both ISRO Ka Band High Throughput Satellite 1 and 2 areidentical.

The first spacecraft shall be built entirely at the contractor’s facility. The assembly,integration and testing of second spacecraft shall be carried out at customer’s premises inIndia. The contractor shall also be responsible to provide all the onboard hardware, softwareand all related accessories and system required for the built of spacecraft in India. It shall bethe responsibility of contractor to get the second spacecraft fabricated in India using themanpower and infrastructure facilities available with ISRO.

The first spacecraft shall be ready for shipment not later than 36 months from the date ofsigning the contract. The second spacecraft shall be ready 6 months after the readiness of firstspacecraft. The date of FMCR completion shall be taken as the readiness date.

Selection of launcher shall be responsibility of customer and contractor shall provide all theinputs required for coordination with launcher agency. It shall be responsibility of contractorto provide all the technical support including shipment of spacecraft to launch base, and AITactivities at launch base to fulfill launch base requirements.

From the launch date, the contractor shall have a maximum of six months to complete thepost launch operations including LEOP operations, performance verification and in orbit testson the spacecraft. The spacecraft operational life of 15 years shall be counted from the datethe spacecraft is handed over to customer after completion of all the above mentionedactivities.

The second spacecraft activities shall start simultaneously with the first spacecraft activitiesand shall follow the same cycle.

The Spacecraft has to meet all requirements as per the Schedules (A to G) when operated atan orbital location TBD deg E in the geostationary orbital arc. Antenna diagrams shall be


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optimized assuming a nominal position of TBD deg E. The exact orbital slot ingeostationary arc of 48 deg E to 112 deg E will be notified.

The Spacecrafts shall be designed to operate with up to eight other collocatedcommunications satellites within the Customer fleet over the Indian Ocean region. TheSpacecraft design shall incorporate all required filtering in its uplink and downlink topreclude interference to and from adjacent satellites and other ITU coordinated RFtransmissions.

Frequency co-ordination will be carried out by ISRO and exact frequencies will be providedduring signing of contract.

2.1 DEFINITIONS2.1.1 SPACECRAFTSpacecrafts hereafter refers to ISRO Ka Band High Throughput Satellite (HTS-1 and HTS-2).

2.1.2 COMPATIBLE LAUNCH VEHICLEMeans the launch vehicles specified in Section 2.1 of Schedule E.

2.1.3 OPERATIONAL LIFEContinuous uninterrupted period starting with the completion of the In Orbit Commissioningof the Spacecraft and during which the Spacecraft operates in orbit fully compliant with thecontractual specifications.

2.1.4 DESIGN LIFEContinuous uninterrupted period during which the Spacecraft and all its subsystems aredesigned to operate in orbit fully compliant with the contractual specifications, notaccounting for life limitation due to propellant depletion.

2.1.5 PROPELLANT LIFEMaximum period during which the Spacecraft can be maintained within its orbital andattitude slots, while fulfilling attitude and orbit control requirements specified in thisdocument and excluding the propellant required to fulfill all activities carried out up to thecompletion of in-orbit commissioning, and comply with residuals, de-orbiting and relocationrequirements.

2.1.6 BEGINNING OF LIFE (BOL)Date corresponding to the completion of In Orbit Commissioning and handing over of theSpacecraft to the customer.

2.1.7 END OF LIFE (EOL)Date corresponding to the end of the Operational Life of the Spacecraft.


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3 RESPONSIBILITYThe contractor is fully responsible for the programme management tasks, as well as reportingto customer on the status of the programme.

3.1 ISRO Ka HTS-1For HTS-1 the contractor is fully responsible for all tasks related to the design, systemengineering, trade-off studies, development, manufacturing, assembly, integration, testing(including testing at the launch site), delivery, storage (when required), launch baseoperations including propellant loading and combined operations, LEOP operations,initialization and deployments, IOT and hot-line support to customer during operationalphase of the spacecraft built according to Schedules A to F, tested according to Schedule Cand meeting the technical specification requirements as defined in Schedule B.The satellite shall meet all the requirements and be capable of being operated during aminimum period of 15 years and shall be compatible with the specified launch vehicles.

3.2 ISRO Ka HTS-2For HTS-2 the responsibility of contractor shall be same as above except that all the AITshall be carried out at customer’s premises by customer manpower under overall control ofcontractor. The facilities & infrastructure which are available at customer premises can beused for the purpose however the contractor has to specify the facilities required.

3.3 ISRO Ka HTS-1 and ISRO Ka HTS-2The satellites shall be made ready for launch with the consumable items (propellants andgases) that are required to fulfill the mission. The satellites shall be made available for theflight model completion review according to the following schedule

a) when the Flight Model Completion Review (FMCR) has been successfullycompleted, the satellite flight model will be considered acceptable for shipment tothe launch site or for storage, if so decided by ISRO;

b) Contractor is fully responsible for transportation of the satellite flight model, allconsumable items (propellants and gases) and support equipment to the launchsite(s);

c) When the Flight Readiness Review (FRR) has been successfully completed, thesatellite flight model will be considered acceptable for launch, with exception of thelaunch preparation activities in the hazardous processing facility and the integratedactivities with the launch vehicle.

Customer shall coordinate the interface with the launch services contractor. Contractor shallensure that the data/documentation of the satellite is available to the launcher agency to carryout associated analysis with the customer chosen launcher.


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3.4 RISK MANAGEMENTContractor shall propose risk management program in order to systematically identify,analyze and handle risks before they become actual problems. As a minimum, risk issues ofreliability, reducibility, design, work to be accomplished, trade-offs, cost, schedules, andother special considerations will be addressed on a regular basis throughout the program andreported to customer. Emphasis shall be put on quantitative, rather than qualitative riskanalysis techniques.

3.5 ANOMALIES HANDLING AND DOCUMENTATIONContractor shall report to customer all anomalies related to hardware similar or identical tothe one used on the spacecraft, observed in other satellites manufactured by contractor, eitherduring ground processing or in orbit, and anomalies reported as industry alerts, which mayaffect the spacecraft. Contractor shall report to customer on such anomalies within one weekor earlier of their occurrences (not necessarily identifying the satellite on which the anomalyoccurred). Contractor shall carry out the root cause analysis and provide the detailed report tothe customer. Also contractor shall be responsible for taking all necessary steps to ensure thatall such deficiencies noted on other satellites manufactured by contractor and deficienciesstemming from industry alerts will be corrected in the spacecraft prior to launch. Similaralerts are also required to be provided during the in-orbit operations of the spacecraft.Contractor shall be responsible for taking all necessary steps to ensure that the deliveredsoftware will reflect the most up to date version of the product line software.

Contractor's responsibilities also include the documentation of all work conducted under theagreement and reporting to the customer on the status and progress of the program asspecified in this schedule.

3.6 EXPORT AUTHORIZATIONSContractor shall be responsible for obtaining all export licenses, export documentation andauthorizations required for the delivery of the items covered by this Statement of Work.Contractor shall report weekly to Customer the progress in achieving the proper licenses, andapply diligence to ensure that all licenses are obtained timely.

4 REALIZATION PLAN & CUSTOMER’S TECHNICALVISIBILITY AND ACCESS

4.1 REALIZATION PLANThe contractor shall provide a detailed realization plan with mandatory inspection / reviewpoints. The same shall be discussed and agreed mutually.


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4.2 TIMELINE AND PROCUREMENTManufacturer shall specify the time line of the contract execution cycle e.g. translation ofmission requirements to system/subsystem requirements/specifications, PDR, CDR,Integrated Test Plan, TRR and any other milestone as perceived fit by both contractor andCustomer etc. Customer’s participation and approval is mandatory during major milestonesevents.

Upon award of the contract, Contractor shall initiate the procurement of all necessary longlead items for both the spacecrafts.

Existing satellite product line resources may be utilized to the extent applicable for theSpacecraft design, design review contents, design review data packages, test plans, testprocedures, analyses and other program documentation.

4.3 ISRO PARTICIPATIONThe active and full participation of ISRO teams during the total execution cycle is asignificant part of the whole process. ISRO requires full involvement during all phases ofSpacecraft build like System Engineering, Configuration finalization, Design, Development,Fabrication and AIT. The activities like overall configuration finalization, Design drivers,System level definitions and Configuration, build plan and process shall be well coordinatedand understood by ISRO team . ISRO engineers are to be involved in all subsystem & systemlevel tests, qualification, acceptance and AIT activities. This may also include somediscussions and knowledge sharing on latest technology domains. Full time participation ofengineers from ISRO is required during the entire spacecraft realization period. ISROengineers will be stationed at Contractor’s premises and shall have full access to ongoingactivities. It is expected that a typical of 25 numbers of ISRO engineers shall be stationed on-site. The contractor shall clearly express their willingness on the subject. Contractor shallallow unimpeded and continuous access by Customer and its designated representatives to theWork carried out in connection with this Agreement by Contractor and its subcontractors atall levels.

For second spacecraft the overall responsibility lies with the contractor and all the AITactivities shall be carried out by customer’s manpower under the overall control andsupervision of contractor at customer’s facility

4.4 HERITAGEThe heritage of systems used onboard the spacecraft are of prime importance to customer. Adetailed heritage matrix of all the systems used onboard the spacecraft shall be provided bythe contractor.


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4.5 ACCESS TO WORKPLACEIn the case, that access by Customer to the Work in progress is impeded by Governmentalrestrictions, Contractor may hire independent experts, with proper Governmental clearance,to oversee the Work on behalf of Customer and to report the progress and findings toCustomer and to Contractor.

Customer shall have approval rights on the selection of the experts.

4.6 SPACECRAFT LIFE AND MASS4.6.1 SPACECRAFT LIFEThe Spacecraft Design Life shall be no less than 18 years. The Spacecraft Operational Lifeshall be at least 15 years.

The Spacecraft design shall be qualified to carry and consume the full capability of itspropellant and pressurant, even if it is not required to meet the spacecraft Life requirements.Upon request by Customer, Contractor shall provide such propellant and pressurant up to themaximum capacity of the tanks at no extra cost, even if the spacecraft Life requirements canbe met with less than full tanks. The decision on loading will be taken during launchcampaign.

4.6.2 SPACECRAFT MASSThe Contractor is expected to propose an optimal dry mass of the spacecraft which meets allfunctional requirements and mission life with 3σ performance. The mass estimate shall bejustified with proper explanations. The contractor shall also keep in view the constraint posedby launch vehicles specified in Schedule E. Since the spacecraft lift-off mass has a directbearing on the launch costs, due weightage will be considered for mass optimized spacecraftconfiguration. A detailed mass budget shall be provided by the contractor.

5 HARDWARE AND SOFTWARE FOR GROUNDSYSTEM VALIDATION TESTING AND SATELLITEOPERATIONS

Ground System Validation Testing to ensure full compatibility between the satellite and theground system will be carried out by the Contractor. The test objectives are to prove thefollowing:

a) Satellite to ground station RF compatibility;b) Satellite to ground system data compatibility;c) Validation of the entire control chain from control centre consoles through to

satellite behavior.

To facilitate satellite operations, tools are required for:


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a) Payload reconfiguration;b) Pointing and rotating the beam of steerable spot antenna;c) In-orbit re-programming of the ADCS subsystem (as applicable).d) Communication Performance Tools

The software for these satellite operations tools shall be provided typically after twelve (12)months from EDC or mutually agreed date from both parties. The software shall be writtenin mutually acceptable programme language. The delivered software shall be portable interms of compatibility with different hardware platforms and operating systems, particularlyLINUX and UNIX.

For all satellite operational software provided by the Contractor, the Contractor shall providefull supporting documentation, build procedures and maintenance tools. Telecommanddatabases, Telemetry database and any other onboard uplinked databases as accepted byISRO shall be provided in electronic format.

5.1 TCR SUITCASEThe Contractor shall deliver a Telemetry, Command and Ranging (TCR) suitcase model to alocation defined by ISRO typically after twelve (12) months from EDC or mutually agreeddate. The suitcase model shall comply with the requirements as given in Schedule B Section2.5.

The Contractor shall be responsible for the transport and insurance of the suitcase to a sitedecided by the Customer. The Contractor shall also be responsible for the transport andinsurance of the suitcase as required forT the validation of the LEOP station network.

The Contractor shall provide a minimum of two days training on the use of the suitcase andshall support the TCR compatibility tests at a site to be specified by the Customer.

The Contractor shall fix any failure or anomaly of the suitcase in less than two (2) weeks.

5.2 TELEMETRY DATAThe Contractor shall deliver Telemetry data for the validation of the ground softwaretypically twelve (12) months from EDC or mutually agreed date.

This data shall be as representative as possible of the in-orbit telemetry data stream. Thecontents and medium shall be agreed with ISRO.

5.3 SATELLITE SIMULATORThe Contractor shall deliver a satellite simulator at the Customer premises or to otherpremises designated by the Customer (i.e. software, hardware as applicable, database andassociated documentation) in accordance with the requirements in Schedule F typically after


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twelve (12) months from EDC or mutually agreed date. The simulator delivered at that timemay include a preliminary launch injection orbit that will be updated when ISRO has decidedon the orbit actually used.

The interfaces between the simulator and the satellite control centre shall be according to theTCP/IP protocols specified by the Customer.

In addition the Contractor shall update the simulator software and database to keep themcompatible with the Operations Handbook and specific satellite flight model data.

5.4 PAYLOAD RECONFIGURATION SOFTWARE TOOLThe Contractor shall provide the necessary software tool to be used by the ground operatorsfor payload reconfiguration, when redundancy and channel selectivity is involved. Thesoftware shall cover any combination of switch, EPC and TWTA failures and include thecapability to define priority channels which should not be disturbed. As a minimum, the newring configuration and the telecommands to be sent shall be provided as an output.

5.5 STEERABLE ANTENNA CONTROL SOFTWARE TOOLThe Contractor shall provide, as appropriate, the necessary software tool to be used by theground operators for the pointing and rotating of the steerable spot antennas. The softwareshall cover the operation and correct pointing of the antennas.

5.6 ADCS RE-PROGRAMMING TOOLFor the cases where ADCS modes need to be re-programmed in orbit, the Contractor shallprovide the relevant command file in a computer medium and a format to be agreed with theCustomer.

5.7 TRAININGThe Contractor shall provide and present to the customer, two separate training courseswithin mutually agreed time. These courses shall enable ISRO staff to maintain the hardwareand software associated with these items, and shall familiarize customer with their detaileddesign.

In addition, the Contractor shall provide and present at the Customer premises two separatetraining courses in order to familiarize the satellite operations personnel with the userinterface and the operational use of these items.

5.8 WARRANTY AND MAINTENANCE POST ACCEPTANCEUnless otherwise agreed by ISRO, the Contractor shall install hardware and software releasesrelated to the items at the customer premises or at any other Indian location specified bycustomer.


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The hardware, software and databases associated with these items shall reflect each satellitedesign and performance at the time of the item release.

The Contractor shall correct any errors/failures/deficiencies detected in the hardware orsoftware associated with each of these items, for the whole satellite lifetime.

During this period, the Contractor shall also update the simulator software and databases toreflect the actual performance of the satellite.

The contractor shall be available for immediate online assistance in case it is required. TheContractor shall position an engineer at the customer’s premises to respond to any customerrequest for an urgent error correction of any of these items within 7 working days. However,customer reserves the right to perform corrections of an urgent nature that will then becommunicated to the Contractor.

The system for submitting, documenting and controlling problems after the initial releaseshall be approved by customer.

5.9 TELECOMMAND ENCRYPTIONIn the event that customer selects the telecommand encryption option, the Contractor shall beresponsible for the security of the flight decryption keys and the surveillance and protectionof such keys including when mounted to the satellite for test or final installation.

6 LAUNCH VEHICLE COMPATIBILITYThe ISRO Ka Band HTS shall be compatible with the candidate launch vehicles as identifiedin Schedule E, Section 2.1.

7 GROUND CONTROL COMPATIBILITYThe Spacecraft shall be controlled from one of the Customer satellite operations facilitieslocated in India. The requirements for Ground Control System are specified in Schedule F.

8 PRODUCT ASSURANCE REQUIREMENTSThe Contractor shall establish and implement a Product Assurance programme in accordancewith the requirements defined in the Product Assurance Requirements (Schedule D).

A Product Assurance Plan shall be prepared by the Contractor describing the programmeactivities and the policies and procedures which will be implemented to ensure thatprogramme objectives are successfully met and addressing the principal areas of customerinvolvement.


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9 SATELLITE REVIEWS AND SCHEDULE9.1 EQUIPMENT QUALIFICATION STATUS REVIEW (EQSR)The purpose of the EQSR is to allow Customer and the Contractor to formally concur on theequipment qualification status. The EQSR shall take place within three (3) months after theEffective Date of Contract (EDC).

9.2 DESIGN REVIEWSThe programme shall incorporate two types of design reviews:a) Preliminary Design Review (PDR);b) Critical Design Review (CDR).

Each design review held for a particular system, subsystem or equipment shall represent thecompletion of a stage in the overall development and qualification process of that system,subsystems or equipment and shall act as decision milestone before proceeding to the nextphase. Detailed definition and objective of these reviews are given in the Assembly,integration and Test plans Requirements (Schedule C).

9.2.1 PRELIMINARY DESIGN REVIEWS (PDR)The objective of the PDR is to review the conceptual design of the satellite, subsystem orequipment and its design and development programme.The completion of the PDR permits the detailed design of the satellite and its subsystems andequipment to proceed.The system PDR shall be conducted within 5 months after the Effective Date of Contract(EDC).

9.2.2 CRITICAL DESIGN REVIEWS (CDR)The objective of the CDR is to provide customer with the opportunity to formally review andto establish confidence in the final design of the satellite, subsystems or equipment. Thecompletion of the CDR permits the Design Standards of the satellite, its subsystems andequipment to be approved. In particular, testing of equipment Engineering QualificationModels and/or Qualification Model will have been completed. The system CDR shall beconducted within 13 months after the Effective Date of Contract (EDC).

9.3 MANUFACTURING, ASSEMBLY, INTEGRATION AND TESTREVIEWS

The manufacturing, Assembly, Integration and Test programme shall incorporate three typesof major review:

a) Satellite Qualification Status Review (SQSR) / Environmental Test Readiness Review(ETRR);

b) Flight Model Completion Review (FMCR);c) Test Reviews.


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A detailed AIT plan to be submitted by contractor which shall be implemented subjected tocustomers approval.

9.3.1 SATELLITE QUALIFICATION STATUS REVIEW (SQSR) /ENVIRONMENTAL TEST READINESS REVIEW (ETRR)

The objectives of this review are to formally review the qualification status of the equipmentand the performance of the satellite and its subsystems and equipment prior to theenvironmental qualification tests and to verify the readiness of the satellite and test facilitiesfor commencement of the environmental testing.

The Satellite Qualification Status Review (SQSR)/(ETRR) of the satellite shall be held priorto the environmental qualification tests on the satellite.

9.3.2 FLIGHT MODEL COMPLETION REVIEW (FMCR)Flight Model Completion Reviews (FMCR) shall be held for all flight model satellites,including a Protoflight Model, at the end of the assembly, integration and test phase of thesatellite.

The successful completion of the Flight Model Completion Review will permit the satellite tobe shipped to the launch site or to be put into storage, if so decided by Customer.

9.3.3 TEST REVIEWSTest Readiness Reviews (TRR) and Test Review Boards (TRB) shall be held respectivelyprior to, and immediately after, major test activities on all equipment, subsystem and satellitemodels (including development models).

The objective of the reviews is to provide ISRO with the opportunity to formally assess thepreparation for the major test activity or the results of the completed tests. These reviews willalso address the test facility and equipments identified for the tests.

9.3.4 IN ORBIT ACCEPTANCE REVIEW (IOAR)The objective of the IOAR is to provide formal acceptance of the satellite in orbit. Thereview will be held after completion of the In Orbit Tests (IOT) and will assess whether thesatellite adequately demonstrates the ability to meet the satellite Technical SpecificationRequirements (Schedule B) over lifetime, taking into account:

a) IOT results for the satellite payload and platform;b) Performance presented at FMCR;c) Overall satellite in flight performance including anomalies experienced and their

resolution;d) Completeness of supporting documentation including End Item Datapacks (EIDPs) and

documentation required for satellite operations;


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10 NON DELIVERABLE EQUIPMENTDevelopment models of equipment shall be retained by the Contractor and the Subcontractorsuntil twelve (12) months after the launch of the last satellite ordered by ISRO in order toassist in the investigation of possible in-orbit anomalies and to evaluate possible remedies.

11 SPARE PROVISIONSAlthough spare provisions (e.g. parts, integrated boards, equipment) are not deliverable items,the Contractor shall ensure that flight spare provisions are available in sufficient number atthe appropriate assembly level to support the tests and pre-launch operations of each satellite,in order to protect the programme schedule in case of failure during ground operations andtesting.

Spare provisions shall fulfill all the applicable specifications of the Contract and shall be fullyinterchangeable with the related flight hardware.

The Contractor's spares philosophy and spares support plan shall be defined and the SpareProvisions shall be identified in the Contractors Hardware Matrix.

12 SATELLITE STORAGE AND SAFE KEEPINGIn case, customer desires to delay the launch, the Contractor shall put a satellite into storagefollowing successful completion of the FMCR. Storage of the satellite will not be longer than2 years from the FMCR of the satellite.

On the request of customer, the Contractor shall support the short term safe-keeping of thesatellite at either the Contractor's site, or if required by customer at the launch site.

13 SERVICESIn addition to the work related to the management of the programme and the design,development, manufacturing, integration, testing, delivery and preparation for launch of thesatellites and associated equipment, the Contractor shall furnish the services describedhereafter.

The detailed requirements for launch, mission support and the LEOP services are specifiedrespectively in Schedule E & F.

13.1 LAUNCH SUPPORT SERVICESThe Launch Support Services include:a) Support for launch vehicle interface co-ordination;b) Final integration, testing, and preparation for launch of satellite at its launch site;


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c) Launch site operations.

13.2 MISSION SUPPORT SERVICESThe Operation Support Services include:a) The preparation of satellite operational documentation for all phases of the satellite

mission (e.g. launch, LEOP, on-station, orbit re-location re-orbiting at the end of life).b) The preparation of in-orbit test plans for the platform and Communication Payload.c) Specialist engineers on site in the on-station control centre or at the Contractor's

premises for detailed investigations concerning on-board anomalies;d) The availability of a facility at the Contractor's premises to monitor the satellite

telemetry of the in-orbit satellites, for performance evaluation and anomalyinvestigation.


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SECTION III

PROGRAM MANAGEMENT AND REALIZATION MILESTONES


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1 PROGRAM MANAGEMENTCustomer considers on time delivery of spacecraft a prime factor of contractor performance,support and evaluation. To that end, contractor should propose the Program Managementplan which shall include but not limited to the following:

Detailed program schedule information flow and management Detailed risk assessment, management and mitigation plan

2 MAJOR REALIZATION MILESTONESCustomer proposes following tentative major realization milestones as given in table S3-2.1.Contractor has freedom to use them as reference and generate their own detailed ISRO Kaband HTS realization plan which will be finalized after mutual agreement.

Table S3-2.1: ISRO Ka-band HTS Realization Milestones

S.N. Milestone Description Date Type

1. Contract Signing Data

2. Project Kick Off Meeting and Systems RequirementReview (SRR)

Review

3. Equipment Qualification Status Review (EQSR) DesignReview

4. Subsystem & Spacecraft Preliminary Design Review(PDR)

DesignReview

5. Issue of payload and platform system specifications Data

6. Subsystem & Spacecraft Critical Design Review (CDR) DesignReview

7. Subsystem Test Readiness Review (TRR) DesignReview

8. Spacecraft level Test Readiness Review (TRR) and TestResults Review (TRR) during different IST phases

DesignReview

9. Complete Spacecraft Level Final Acceptance Test priorto Launch Site shipment

Delivery

10. Spacecraft Pre Shipment Review (PSR) or Flight ModelCompletion Review (FMCR)

DesignReview

11. Spacecraft shipment and Pre-launch Test Result Review Delivery

12. Spacecraft In Orbit Acceptance Review (IOAR) DesignReview

ISRO


ISRO Ka BAND HTS

SCHEDULE B

TECHNICAL SPECIFICATIONSREQUIREMENT

ISRO/ISAC/HTS/RFP/001/06VER-01 SCHEDULE - B

TECHNICAL SPECIFICATIONS REQUIREMENT

December 2015 Page B-3ISRO CONFIDENTIAL PROPRIETARY

TABLE OF CONTENTS

SCOPE ........................................................................................................................................9

1 COMMUNICATION PAYLOAD REQUIREMENTS ...........................................................9

1.1 INTRODUCTION ..............................................................................................................9

1.1.1 SPECTRUM ALLOCATION................................................................................................9

1.1.2 COVERAGE ...................................................................................................................10

1.1.3 BEAM LAYOUT .............................................................................................................12

1.1.4 PAYLOAD CONFIGURATION .........................................................................................14

1.1.5 PAYLOAD PERFORMANCE SPECIFICATIONS .................................................................15

1.1.6 KA-BAND BEACON SPECIFICATIONS ............................................................................30

1.1.7 SPECIFICATIONS OF IOT-HORN .....................................................................................31

1.1.8 OPTIONAL CONFIGURATION (INDIA BEAMS + STEERABLE BEAMS) .............................31

2 TELEMETRY, COMMAND AND RANGING (TCR) SYSTEM ..........................................33

2.1 GENERAL REQUIREMENTS............................................................................................33

2.2 COMMAND SUBSYSTEM ...............................................................................................34

2.2.1 GENERAL REQUIREMENTS ...........................................................................................34

2.2.2 CONFIGURATION ..........................................................................................................35

2.2.3 FREQUENCY OF OPERATION.........................................................................................36

2.2.4 FREQUENCY STABILITY ...............................................................................................36

2.2.5 ANTENNA COVERAGE ..................................................................................................36

2.2.6 ANTENNA POLARIZATION ............................................................................................36

2.2.7 ANTENNA AXIAL RATIO ...............................................................................................37

2.2.8 INPUT POWER FLUX DENSITY....................................................................................... 37

2.2.9 MODULATION ...............................................................................................................37

2.2.10 BASEBAND REQUIREMENT ...........................................................................................37

2.2.11 COMMAND ENCRYPTION (OPTIONAL) ..........................................................................38

2.3 TELEMETRY SUBSYSTEM .............................................................................................39


2.3.2 CONFIGURATION ..........................................................................................................39

2.3.3 FREQUENCY OF OPERATION.........................................................................................40

2.3.4 FREQUENCY STABILITY ...............................................................................................40




2.3.5 ANTENNA COVERAGE ..................................................................................................41

2.3.6 ANTENNA POLARIZATION ............................................................................................41

2.3.7 ANTENNA AXIAL RATIO ...............................................................................................41

2.3.8 EFFECTIVE ISOTROPIC RADIATED POWER (EIRP) .........................................................41

2.3.9 EIRP STABILITY ............................................................................................................41

2.3.10 MODULATION ...............................................................................................................42

2.3.11 BASE-BAND REQUIREMENTS ........................................................................................42

2.4 RANGING SUBSYSTEM ..................................................................................................42


2.4.2 RANGE TONE FREQUENCIES AND MODULATION .........................................................43

2.4.3 RANGE CALIBRATION...................................................................................................43

2.4.4 MODULATION ...............................................................................................................43

2.4.5 RANGING PERFORMANCE REQUIREMENT ....................................................................43

2.5 GROUND STATION COMPATIBILITY UNIT (SUITCASE MODEL) ......................................44

3 ELECTRICAL POWER SYSTEM.................................................................................... 44

3.1 GENERAL REQUIREMENTS............................................................................................44

3.2 SOLAR ARRAY ..............................................................................................................45

3.3 BATTERIES....................................................................................................................47

3.4 POWER CONTROL & PROTECTIONS ...............................................................................49

3.5 REDUNDANCY ..............................................................................................................51

3.6 POWER BUDGET............................................................................................................52

4 THERMAL REQUIREMENTS ........................................................................................53

4.1 THERMAL RANGES AND TEMPERATURES.....................................................................53

4.2 THERMAL DESIGN.........................................................................................................54

4.3 THERMAL MATHEMATICAL MODELS AND ASSOCIATED DOCUMENTATION ................56

4.4 THERMAL SUBSYSTEM DEVELOPMENT TESTING .........................................................58

4.4.1 ANALYSIS / MEASUREMENT .........................................................................................58

4.4.2 HEAT PIPE / LOOP HEAT PIPE DEVELOPMENT TESTS: ...................................................59

4.4.3 DEPLOYED THERMAL RADIATOR .................................................................................59

4.4.4 SPACECRAFT THERMAL BALANCE TEST: .....................................................................59

4.5 THERMAL CONTROL MATERIAL AND PROCESSES ........................................................59

5 STRUCTURE SUBSYSTEM ............................................................................................60




5.1 FUNCTIONAL REQUIREMENTS: .....................................................................................60

5.2 PERFORMANCE REQUIREMENTS:..................................................................................60

5.3 DESIGN REQUIREMENTS: ..............................................................................................60

5.3.1 ENVIRONMENTAL LOADS .............................................................................................60

5.3.2 DESIGN CRITERIA .........................................................................................................61

6 ATTITUDE DETERMINATION AND CONTROL SYSTEM(ADCS) ..................................63

6.1 DEFINITION OF SPACECRAFT AXES............................................................................... 63

6.2 ATTITUDE CONTROL SUBSYSTEM FUNCTIONAL REQUIREMENTS............................... 63

6.3 ATTITUDE DETERMINATION .........................................................................................63

6.4 ATTITUDE CONTROL.....................................................................................................64

6.4.1 STABILITY ....................................................................................................................64

6.4.2 POINTING ACCURACIES ................................................................................................65

6.4.3 STATION KEEPING MANEUVERS:..................................................................................66

6.4.4 IN-ORBIT ANTENNA PATTERN MEASUREMENT: ...........................................................66

6.4.5 CONTROL BIAS CAPABILITY: ........................................................................................66

6.4.6 SAFE MODES AND RE-ACQUISITION: ............................................................................67

6.4.7 PAYLOAD SHEDDING IN EMERGENCY: .........................................................................68

6.4.8 MASS PROPERTIES ........................................................................................................68

6.4.9 AUTONOMY ..................................................................................................................69

6.4.10 CONTROL ELECTRONICS FAULT PROTECTION .............................................................69

6.4.11 PROPULSION INTERFACES ............................................................................................69

6.4.12 TELEMETRY INTERFACES.............................................................................................69

6.4.13 SPECIFIC ADCS TEST:....................................................................................................70

6.4.14 SATELLITE ON BOARD PROCESSOR MEMORY DATA ....................................................70

6.4.15 ADCS BLOCK DIAGRAMS ..............................................................................................70

6.4.16 ATTITUDE CONTROL AND DETERMINATION DATA ......................................................70

6.4.17 FUEL SLOSHING DATA..................................................................................................70

7 PROPULSION SYSTEM .................................................................................................71

7.1 SUBSYSTEM REQUIREMENTS .......................................................................................71

7.2 OPERATIONAL REQUIREMENTS....................................................................................71

7.2.1 RELIABILITY.................................................................................................................71

7.2.2 SAFETY .........................................................................................................................71




7.2.3 OPERATIONAL LIFETIME ..............................................................................................71

7.2.4 STORAGE LIFE ..............................................................................................................72

7.2.5 MAINTAINABILITY .......................................................................................................72

7.2.6 GROUND OPERATIONS..................................................................................................72

7.2.7 LAUNCH........................................................................................................................72

7.2.8 TRANSPORTABILITY .....................................................................................................72

7.2.9 FILLING AND DRAINING ...............................................................................................72

7.3 DESIGN AND CONSTRUCTION .......................................................................................73

7.3.1 PARTS, MATERIALS AND PROCESSES ...........................................................................73

7.3.2 MECHANICAL DESIGN ..................................................................................................73

7.3.3 THERMAL DESIGN DESCRIPTION ..................................................................................73

8 MECHANISMS AND PYROTECHNICS ..........................................................................73

8.1 DESIGN REQUIREMENTS ...............................................................................................73

8.2 PYROTECHNIC DEVICES ...............................................................................................75

9 ORBITAL OPERATIONS ...............................................................................................76

9.1 TRANSFER ORBIT ..........................................................................................................76

9.1.1 ATTITUDE ERROR BUDGET...........................................................................................76

9.1.2 TRACKING NETWORK ...................................................................................................77

9.2 STATION-KEEPING ........................................................................................................77

9.2.1 TERMINOLOGY .............................................................................................................77

9.2.2 REQUIREMENTS ............................................................................................................78

9.2.3 DRIFT AND ECCENTRICITY CONTROL MANEUVERS .....................................................79

9.2.4 INCLINATION CONTROL MANEUVERS..........................................................................79

9.2.5 HYBRID INCLINATION/ECCENTRICITY CONTROL MANEUVERS ...................................79

9.3 ADDITIONAL REQUIREMENTS ......................................................................................79

9.3.1 SPACECRAFT LONGITUDE RELOCATION ......................................................................79

9.3.2 RANGING SYSTEM REQUIREMENTS.............................................................................. 79

9.3.3 MAXIMUM DELTA-V FOR ATTITUDE RECOVERY MANEUVERS ....................................79

9.3.4 DISPOSAL ORBIT AND PASSIVATION ............................................................................80

9.4 PROPELLANT BUDGET ..................................................................................................80

10 DESIGN AND ANALYSES ..............................................................................................81

10.1 DESIGN ANALYSES AND STUDY REPORTS ....................................................................81




10.1.1 REQUIREMENTS: ...........................................................................................................81

10.1.2 ANALYSES DATA ..........................................................................................................82

10.2 SYSTEM LEVEL ANALYSES ...........................................................................................82

10.2.1 DYNAMIC ANALYSES ...................................................................................................82

10.2.2 PROPELLANT BUDGET ANALYSES................................................................................82

10.2.3 MISSION ANALYSES......................................................................................................83

10.2.4 MASS PROPERTIES ANALYSES ......................................................................................83

10.2.5 ELECTROMAGNETIC COMPATIBILITY (EMC) ANALYSES..............................................83

10.2.6 ENVIRONMENTAL EFFECTS ANALYSES ........................................................................83

10.2.7 PAYLOAD ANALYSES ...................................................................................................84

10.2.8 FDIR RESPONSE AND OBSERVABILITY .........................................................................85

10.3 SUBSYSTEM LEVEL ANALYSES..................................................................................... 85

10.3.1 TELEMETRY, TELECOMMAND AND RANGING ANALYSES............................................ 85

10.3.2 THERMAL ANALYSES ...................................................................................................85

10.3.3 STRUCTURAL ANALYSES .............................................................................................86

10.3.4 ELECTRICAL POWER ANALYSES ..................................................................................87

10.3.5 ATTITUDE DETERMINATION AND CONTROL ANALYSES ..............................................88

10.3.6 PROPULSION ANALYSES ...............................................................................................88




SCOPEThis document defines the performance and operational requirements for ISRO Ka BandHigh throughput Satellites and the associated support equipment.

1 COMMUNICATION PAYLOAD REQUIREMENTS1.1 INTRODUCTIONThe Ka-band communication system shall consist of space segment, which includes highthroughput Ka-band communication satellite having multi-beam configuration, and groundsegment which includes hub stations as well as user terminals. Contractor shall propose onlyfor the space segment.

The satellite payload configuration details are described in this section. The primary payloadconfiguration shall have minimum 88 spot beams over the specified coverage. As anadditional option, the contractor shall also propose 4 independently steerable beams.

The communication payload transponders shall have bent-pipe configuration. The satelliteshall provide fixed broadband services to multiple users in star-based network. The forwardlink shall be from hub to user terminal via satellite forward channel. The return link shall befrom user terminal to hub via satellite return channel.

1.1.1 SPECTRUM ALLOCATIONThe communication system shall have uplink and downlink frequencies in Ka-band for bothforward and return links. The polarizations used shall be Circular (LHCP/RHCP) or Linear(LV/LH). There shall be two downlink Ka-band beacons at TBD GHz in two orthogonalpolarizations. There shall be one uplink Ka-band beacon at TBD GHz for RF tracking inTBD polarization. The exact frequencies and polarizations will be notified later.

1.1.1.1 USER SPECTRUMUser spectrum available for uplink will be 500 MHz in both polarizations, out of totalspectrum of 27.0 GHz to 31.0 GHz.

User spectrum available for downlink will be 500 MHz in both polarizations, out of totalspectrum of 17.7 GHz to 21.2 GHz.

The spectrum in forward and in return links shall be available as split bands. The exact 500MHz uplink and downlink user spectrum and polarization will be notified later.

1.1.1.2 HUB SPECTRUMHub spectrum available for uplink will be 1 GHz in both polarizations, out of total spectrumof 27.0 GHz to 31.0 GHz.




Hub spectrum available for downlink will be 1 GHz in both polarizations, out of totalspectrum of 17.7 GHz to 21.2 GHz.

The spectrum in forward and in return links shall be available as split bands. The exact 1 GHzuplink and downlink hub spectrum and the polarization will be notified later.

As 1GHz spectrum is available for hub beams in both polarizations, each hub shall have 2GHz of total spectrum. Thus, each hub can cater to 8 user beams (4 user beams in eachpolarization).

The frequency details are provided in Table 1.1.

Table 1.1: Spectrum Allocation

S.N. Parameter FrequencyRange (GHz)

Bandwidth(MHz)

Polarization

1. Hub uplink 27.0 – 31.0 1000LHCP & RHCP

ORLH & LV

2. Hub downlink 17.7 – 21.2 1000LHCP & RHCP

ORLH & LV

3. User uplink 27.0 – 31.0 500LHCP & RHCP

ORLH & LV

4. User downlink 17.7 – 21.2 500LHCP & RHCP

ORLH & LV

5. Bandwidth per user beam - 250LHCP/RHCP

ORLH/LV

Note to Contractor: [Exact frequencies will be confirmed at the time of contract.]

1.1.2 COVERAGEThe coverage includes Indian mainland and Andaman & Nicobar and Lakshadweep Islands.The tentative coverage polygon is shown in Figure 1.1 and polygon points are provided inTable 1.2. The coverage polygon will be finalized at the time of contract.




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SAT

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40.00

30.00

20.00

10.00

0.00

70.00

Figure 1.1: Coverage Polygon

Table 18.20.:00Polygon Points90.00 100.00East Longi tude (Degrees)

Indian Mainland Polygon Points

S.N. Station Longitude(E)

Latitude(N) Station Longitude

(E)Latitude

(N)

1. 01 72.49 35.93 27 92.62 21.872. 02 73.50 36.87 28 91.92 23.743. 03 75.51 37.15 29 91.86 24.304. 04 77.46 35.53 30 90.31 25.005. 05 79.08 36.09 31 89.68 26.246. 06 81.15 35.31 32 88.68 26.457. 07 79.25 31.35 33 88.68 26.458. 08 80.97 30.34 34 88.68 26.439. 09 80.36 29.00 35 89.00 21.6410. 10 83.96 27.50 36 86.90 20.0011. 11 85.80 26.86 37 80.82 15.2612. 12 87.98 26.34 38 80.03 10.0613. 13 87.85 28.38 39 77.63 7.86




Indian Mainland Polygon Points

S.N. Station Longitude(E)


(E)Latitude

(N)

14. 14 87.84 28.38 40 76.28 8.6315. 15 87.84 28.39 41 73.10 16.2916. 16 89.28 28.09 42 72.04 20.6417. 17 89.28 28.09 43 70.48 20.4718. 18 88.91 27.10 44 68.69 21.8719. 19 88.90 27.09 45 67.52 24.2120. 20 90.98 26.81 46 70.54 24.6621. 21 91.87 28.11 47 68.97 27.0522. 22 95.22 29.68 48 70.20 28.5023. 23 97.46 28.50 49 71.51 28.3924. 24 97.24 27.05 50 74.45 32.3025. 25 95.54 26.66 51 73.50 32.9126. 26 92.62 21.87

S.N.

Andaman & Nicobar IslandsPolygon Points

Lakshadweep IslandsPolygon Points

Station Longitude(E)


(E)Latitude

(N)

1. A01 92.2 11.0 L01 72.11 12.102. A02 92.4 9.7 L02 73.95 11.293. A03 92.8 8.6 L03 73.59 7.724. A04 93.2 7.5 L04 71.82 7.725. A05 93.7 6.4 L05 72.17 12.106. A06 94.3 6.47. A07 94.0 8.48. A08 93.4 9.69. A09 94.5 13.410. A10 92.8 13.8

1.1.3 BEAM LAYOUTThe coverage shall be realized using multiple spot beams. The Ka-band multi beam payloadshall consist of minimum of 88 beams of about 0.4 degrees each. Contractor to provide theantenna gain roll-off contours beyond the coverage polygon up to an extent of 30 dB belowEOC, in steps of 2 dB.




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1.1.3.1 USER BEAMS CONFIGURATIONA typical beam arrangement over the coverage area is shown in Figure 1.2. The beampositions will be finalized at the time of contract to meet beam to beam isolation, coverageand other requirements (including antenna design considerations). The beam layout should beoptimized to have major metro cities near the beam centre. Any other suitable alternatepropsal is acceptable after discussion with customer.

a) Frequency Reuse SchemeThe user beam configuration shall have four color re-use plan. The total available userspectrum is 500 MHz in both polarizations (i.e. total spectrum of 1 GHz). Each user beam isalloted 250 MHz in one of the polarizations.

The frequency and polarization of a particular beam shall depend on the final beam layoutand the re-use plan. The beam layout in Figure 1.2 is representative and should be optimizedby contractor for best performance and capacity. The same will be finalized at the time ofcontract.

40.00

30.00

20.00

10.00

60.00 70.00 80.00 90.00 100.00East Longitude (Degrees)

Figure 1.2: Typical Beam Arrangement




1.1.3.2 HUB BEAMS CONFIGURATIONAll the hubs will be located within Indian mainland. Contractor should optimize the hublocations for optimum performance (proximity with large cities is desirable). Contractorshould propose a redundancy scheme for the hub stations. Hub locations as well as beamlayout shall be finalized at the time of contract.

1.1.4 PAYLOAD CONFIGURATIONContractor to provide mass, power and thermal dissipation budgets for the proposedcommunication payload configuration. There shall be Ka-band horns for beacon operations &In-Orbit Test.

Contractor should configure the payload to enable in-orbit characterization of all forward andreturn transponders. The method to characterize transponders during IOT shall part oftechnical proposal, detailing the measurements of major parameters.

1.1.4.1 KA-BAND BEACONThere shall be two Ka-band beacon transmitters (source + SSPA). Each transmitter shall haveunique frequency and comprise of two redundant chains, one beacon source at TBD andTBD GHz and one beacon SSPA for amplification. The amplified beacon signal will bepassed through a harmonic reject filter (HRF) to transmit horn. Each transmitter shall havedifferent polarization. The actual frequency and polarization for both beacons shall befinalized later.

1.1.4.2 ANTENNA RF TRACKING SYSTEMThe antenna pointing requirements of the multi-beam payload shall be met by using anonboard RF tracking system along with a fine pointing and trimming mechanism connectedto each of the reflectors.

In order to meet the payload performance specifications, the RF Tracking system should beable to provide the pointing accuracy better than 0.04 degrees. However, contractor maypropose any better accuracy with improved performance. The ground beacon transmitterfrequency and polarization will be finalized later. Contractor should provide the EIRP andother performance specifications for ground beacon transmitter. The ground beacon locationshall be preferably same as hub locations.

1.1.4.3 PAYLOAD REDUNDANCYThe minimum redundancy scheme for all active subsystems is given below. Contractor mayoptimize the redundancy ring size as per requirement while maintaining the same redundancyratio and without degrading reliability. However, contractor may propose alternate scheme(with adequate justification) to meet the reliability requirement.




a) Ka-Band Transponders:The minimum redundancy for all active subsystems of Ka-band transponders shall be asgiven below:

Low Noise Amplifier (LNA): [30:25]Down-Converter (DoCon): [26:22]High Power Amplifier (HPA): [26:22]

b) Ka-Band Beacon Transmitters:The minimum redundancy for all active subsystems of Ka-band beacon transmitter shall beas given below:

Beacon Source: [2:1]Beacon SSPA: [2:1]

1.1.5 PAYLOAD PERFORMANCE SPECIFICATIONSThe Table 1.4a shows the mandatory performance specifications for forward as well as returnlinks.

Table 1.4b shows the desirable performance specifications for forward as well as return links.Contractor should also supply the expected values for all desirable parameters in addition tomandatory specifications. These parameters may be traded off depending on individualsubsystem designs. In case of any non-conformance in desirable parameters, the trade-offanalysis shall be provided along with technical justification to show that the impact of systemperformance is negligible and acceptable. However, if the justification is found insufficient,ISRO may treat it as Non-Compliance.

Contractor should ensure compatibility to the following signals during operation oftransponder:

a. Single Carrier (CW) at any frequency within any operating uplink band.b. Single Carrier Digital signals with any or combination of ModCods prescribed in

DVB-S, DVB-S2, DVB-RCS2, and DVB-S2x (Extension) standards (e.g. PSK, QAMand APSK)

c. Multiple carriers (MF-TDMA or FDMA)d. Single TDMA carrierse. Combination of any of the above typesf. Any other emerging standard format for HTS applications

Note to Contractor: [The contractor should indicate the limitations in signal compatibility, ifany]

Contractor shall keep the provision for characterization and verification of the applicableperformance parameters, including antenna pattern measurement, for all the forward andreturn links during IOT. There should also be provision to measure the end to endperformance of transponder channels at the spacecraft integrated level. The performancespecifications should be met under all spacecraft environmental conditions.




The description of all the performance parameters is provided after the specification Tables(refer Section 1.1.5.1 to 1.1.5.39).

Table 1.4a: Payload Performance Specifications (Mandatory)

S.N. Performance parameters Units

Specifications

Forward Return

1.EIRP (Single carrier saturation)

- Over 60% coverage area- Over 95% coverage area

dBW>69.5> 66.5

-

2.EIRP (Single carrier saturation)

- Over 60% beam area- Over 95% beam area

dBW - > 69.5> 66.5

3.G/T

- Over 60% coverage area- over 95% coverage area

dB/K -> 21> 18

4.G/T

- Over 60% beam area- over 95% beam area

dB/K> 21> 18

-

5. Bandwidth- Per user beam (either polarization)- Per hub beam (each polarization)

MHz 2501000

2501000

6.Polarization senseReceiveTransmit

-Refer Table 1.1, Page

B 10

7.

Multi-beam System C/I- User beams (Over 60% beam area)- Hub beams (Over Hub locations

and 40% beam area)dB

> 17> 25

> 20> 25

8. Coverage Requirement - Refer Section 1.1.2

9. Transponder Modes of Operation - 1. FGM, 2. ALC,3. Mute

10. Saturation Flux Density Range dBW/m2 -[90+(G/T)] TO-[70+(G/T)]

11. ALC Mode O/P Power Variation dB pp< 1




S.N. Performance parameters Units

Specifications

Forward Return

12.

Noise Power Ratio At Given OBO- 2.5 dB- 3.0 dB- 3.5 dB- 4.0 dB- 6.0 dB

dB

>13>15>17>18>22

>13>15>17>18>22

13.

EIRP Stability (excluding antennaeffects)Over any operating dayOver operating life

dB pp < 1< 1.5

< 1< 1.5

14.

G/T stability (excluding antennaeffects)Over any operating dayOver operating life

dB pp < 1< 1.5

< 1< 1.5

Note: All specifications with geographical constraints (i.e. EIRP, G/T or C/I, over percentageof beam or coverage area) shall be computed using grid based approach. A grid should bedefined over coverage polygon or beam, as applicable, with the grid point spacing less than1/10th of the beam diameter. The parameter values should be derived based on the minimumvalue of the desired parameter over the specified percentage of points.

Table 1.4b: Payload Performance Specifications (Desirable)

S. N. Performance parameters Units

Specifications

Forward Return

1. EIRP (Single carrier saturation) over100% coverage area

dBW 63.5 -

2. G/T over 100% coverage area dB/K - 15

3.

EIRP Stability (including antennaeffects)Over any operating dayOver operating life

dB pp < 2

< 3

< 2

< 3





Specifications

Forward Return

4.

Small signal gain stability(excluding antenna effects)Over any operating dayOver operating life

dB pp< 2< 4

< 2< 4

5.

G/T stability (including antennaeffects)Over any operating dayOver operating life

dB pp < 2< 3

< 2< 3

6. Multi-beam system C/I Over 100%coverage area

dB > 10 > 10

7.

Wideband receive response rejectionat

- 800 MHz from edges of receiveband

- 500 MHz from edges of receiveband

- Transmit frequency bands

dB> 30

-> 130

-> 30>130

8.

Commandable gain in FGM- Gain step range- Gain step size- Tolerance

dB > 301

< ±0.5

> 301

< ±0.5

9.

ALC mode specifications- Dynamic range- Output power control attenuator

(OPCA) range- OPCA step size- Tolerance

dB

dBdBdB

> 30

>150.5

< ±0.5

> 30

>150.5

< ±0.5

10.

FREQUENCY CONVERSIONMHZ TBD

TBDTBDTBDa) Translation frequency*

b) Initial setting accuracy ppm < ± 0.5 < ± 0.5c) Net translation error- Over a day- Over any month which includes

eclipse- Over life time

ppm< ± 1

< ± 2.5

< ± 10

< ± 1< ± 2.5

< ± 10





Specifications

Forward Return

d) Phase noise SSB level atfrequency offset

- 10.0 Hz- 100.0 Hz- 1.0 kHz- 10.0 kHz- 100.0 kHz- 1.0 MHz- 10.0 MHz

dBc/Hz

< -33< -62< -81< -84< -94< -94< -94

< -33< -62< -81< -84< -94< -94< -94

11.

Linearity Of Receive Section: ThirdOrder Imp Level Due To TwoAdjacent Channel Carriers OperatedAt Saturation At Low Gain Setting

dBc< -44 < -44

12.

In-band frequency response(saturation)

- over 40% BW- over 70% BW- over 80% BW- over 100% BW

dB-pp< 1.1< 1.4< 1.7< 2

< 1< 1.5< 1.8< 3.5

13.

In-Band Frequency Response (SmallSignal)

- Over 40% BW- Over 70% BW- Over 80% BW- Over 100% BW

dB/pp< 1.5< 2

< 2.5< 4

< 1.2< 1.8< 2.3< 4

14.

Input Gain Slope Response- Over 70% BW- Over 90% BW- Over 100% BW

dB/MHz< 0.1

< 0.15< 0.2

< 0.1< 0.15< 0.2

15.

Total Gain Slope Response- Over 70% BW- Over 90% BW- Over 100% BW

dB/MHz< 0.2

< 0.25< 0.4

< 0.2< 0.25< 0.4





Specifications

Forward Return

16.

In-Band Group Delay (Input)- Over 40% BW- Over 70% BW- Over 80% BW- Over 90% BW- Over 100% BW

ns

< 5< 10< 15< 25< 30

< 5< 10< 15< 25< 30

17.

In-Band Group Delay (Total)- Over 40% BW- Over 70% BW- Over 80% BW- Over 90% BW- Over 100% BW

ns

< 10< 20< 30< 45< 60

< 10< 20< 30< 45< 60

18.

Receive Narrowband Rejections- CF ± 262.5- CF ± 285- CF ± 500 MHz & beyond- CF ± 137.5- CF ± 160- CF ± 250 MHz & beyond

dB

> 15> 30> 40

---

---

> 15> 30> 40

19.

Transmit narrowband rejections- CF ± 137.5- CF ± 160- CF ± 250 MHz & beyond- CF ± 262.5- CF ± 285- CF ± 500 MHz & beyond

dB

> 10> 25> 40

---

---

> 10> 25> 40

20.

Total channel phase-shift (w.r.t. 20dB IBO) at

0 dB IBO-3 dB IBO-6 dB IBO

deg p-p < 20< 15< 10

< 20< 15< 10





Specifications

Forward Return

21.

AM/PM transfer coefficient at totalflux density relative to SFB

- 0 dB- 3 dB- 6 dB- 9 dB- 12 dB- 14 dB

deg/dB

< 8< 6< 4< 3< 2< 2

< 8< 6< 4< 3< 2< 2

AM/PM conversion coefficient deg/dB < 5 < 5

22.

Spurious output(All levels referred to transmitantenna input)a) Spurious within the transmit band dBc < -60 < -60b) Spurious outside the transmitband, in 4 kHz band.- Transmit harmonics and otherproducts.

dBW< -60 < -60

c) Spurious due to receiver L.O, atmaximum SFD dBc < -80 < -80

23.Overdrive capability level abovenominal operation at maximum inputflux density of channel

dB20 20

* Depends on Uplink & Downlink frequencies

The description of payload performance parameters is given below.

1.1.5.1 EFFECTIVE ISOTROPIC RADIATED POWER (EIRP)The EIRP specification (refer Table 1.4a) shall refer to single carrier saturated TWTA power.The single carrier saturation EIRP is to be calculated using the following formula:Single Carrier Saturated EIRP = HPA O/P Power# + Antenna Gain$ – O/P Losses#HPA output power corresponds to Single carrier Saturated TWTA output power$Antenna Gain for EIRP over 60% coverage should be taken as minimum antenna gain over60% coverage area. Antenna gain for EIRP over 95% coverage should be taken as minimumantenna gain over 95% area covered by the specified coverage polygon. In the case of EIRPin return (User to Hub) links, the EIRP within the required percentage of beam area should beconsidered.




A minimum specified EIRP is to be met over 100% coverage area in Forward link (referTable 1.4b). Contractor should provide the peak EIRP values and ensure that the power fluxdensity limits recommended by ITU are not exceeded.

1.1.5.2 G/TThe specification of G/T shall be as per Table 1.4a. A uniform brightness temperature valueof 290 K shall be assumed for the calculation of antenna G/T performance. Antenna gain forG/T specified over 60% coverage should be taken as minimum receive antenna gain over60% area covered by specified coverage polygon. Antenna gain for EIRP over 95% coverageshould be taken as minimum antenna gain over 95% area covered by the specified coveragepolygon. In the case of G/T in forward (Hub to User) links, the G/T within the requiredpercentage of beam area should be considered.

A minimum specified G/T is to be met over 100% coverage area in Return link (refer Table1.4b). Contractor should also provide the peak G/T values.

1.1.5.3 BANDWIDTHThe bandwidth for user beam as well as hub beam should be as per the specification (referTable 1.4a). This specifies the total bandwidth of the transponder including the guard band.The usable bandwidth of each user beam shall be at least 90% of this value. The hubbandwidth should support four such user beams.

1.1.5.4 POLARIZATIONThe polarization is specified as circular or linear (refer Table 1.4a). The sense of polarizationfor each beam shall depend on the final beam layout and frequency re-use plan.

1.1.5.5 MULTI-BEAM SYSTEM C/IMulti-beam System C/I over at least 60% of user beam area and anywhere over the frequencyband should be as per the specification (refer Table 1.4a). This specification indicates themaximum aggregate interference at the receiver (ground or onboard) from the cross-polarization isolation of same beam as well as interference from other co-polar interferingbeams and cross-polar interfering beams. For the hub beams the specification is over 40% ofthe beam and on the location at which the hubs are suggested by the contractor.

The contractor should provide analytical/simulated & measured C/I plots over entirecoverage area for compliance.

1.1.5.6 COVERAGE REQUIREMENTAll the specifications should be met over the specified coverage area as defined in Section1.1.2.




1.1.5.7 TRANSPONDER MODES OF OPERATIONThe transponder should be capable of operating in Fixed Gain Mode (FGM), AutomaticLevel Control (ALC) mode and Mute Mode. The modes shall be selectable through groundcommand. The description of these modes is given below:

a) Fixed Gain Mode (FGM)In fixed gain mode, the transponder shall provide constant gain over the operating SFDrange, except near HPA saturation (due to gain compression at saturation).

b) Automatic Level Control (ALC)In ALC mode, the output power level of the transponder shall remain constant over aspecified input level range referred to as dynamic range of transponder. For detailedspecifications, refer section 1.1.5.9 and 1.1.5.19.The ALC time constant shall be within 10 to 50 ms.

c) Mute ModeIn Mute mode, the transponder gain shall be reduced by at least 40 dB relative tominimum FGM gain state.When an HPA is powered ON, it is desirable to initiate in Mute mode or in the lowest-gain state available in FGM.

1.1.5.8 SATURATION FLUX DENSITY RANGEThe Saturation Flux Density (SFD) range for all the transponders over the entire coveragearea should be as per the specification (refer Table 1.4a). This SFD should be met in FGM aswell as ALC mode of operation. The G/T to be considered for the SFD specification is theaverage G/T of the beam computed using the grid based approach given above.

1.1.5.9 ALC MODE O/P POWER VARIATIONThis specification indicates the ALC mode output power variation over the ALC dynamicrange (refer Section 1.1.5.19).

1.1.5.10 NOISE POWER RATIO AT GIVEN OBOThe specification for NPR (refer Table 1.4a) describes the nonlinearity of the end to endtransponder from receive antenna feed output to transmit antenna feed input for both forwardand return channel paths.

1.1.5.11 EIRP STABILITY (EXCLUDING ANTENNA EFFECTS)The EIRP in the direction of any location within the service area shall not vary by more thanpeak to peak values (shown in Table 1.4b) over any operating day and over lifetime of thesatellite. At the end of life, the specified EIRP stability shall include all degradation factors.This specification does not include the stability variations due to antenna effects.




1.1.5.12 G/T STABILITY (EXCLUDING ANTENNA EFFECTS)The G/T in the direction of any location within the service area shall not vary by more thanpeak to peak values (shown in Table 1.4b) over any operating day and over lifetime of thesatellite. At the end of life, the specified G/T stability shall include all degradation factors.

This specification does not include the stability variations due to antenna effects.

1.1.5.13 EIRP STABILITY (INCLUDING ANTENNA EFFECTS)The EIRP in the direction of any location within the service area shall not vary by more thanthe peak-to-peak values (shown in Table 1.4b) over any operating day and over lifetime ofthe satellite. At the end of life, the specified EIRP stability shall include all degradationfactors. This specification includes all antenna effects, taking into account the performance ofRF tracking system and antenna pointing mechanism.

1.1.5.14 SMALL SIGNAL GAIN STABILITY (EXCLUDING ANTENNA EFFECTS)The small signal gain of transponder when operating in FGM mode and without the use ofchannel gain control shall not exceed the peak-to-peak values as shown in Table 1.4b.

1.1.5.15 G/T STABILITY (INCLUDING ANTENNA EFFECTS)The G/T in the direction of any location within the service area shall not vary by more thanthe peak-to-peak values shown in Table 1.4b over any operating day and over lifetime of thesatellite. At the end of life, the specified G/T stability shall include all degradation factors.This specification includes all antenna effects, taking into account the performance of RFtracking system and antenna pointing mechanism.

1.1.5.16 MULTI-BEAM SYSTEM C/I (OVER 100% COVERAGE AREA)Multi-beam System C/I within the entire coverage area and anywhere over the frequencyband should be as per the specification (refer Table 1.4b). This specification indicates themaximum aggregate interference at the receiver (ground or onboard) from the cross-polarization isolation of same beam as well as interference from other co-polar interferingbeams and cross-polar interfering beams.

1.1.5.17 WIDE BAND RECEIVE RESPONSE REJECTIONThe specification for wideband receive response shall be as per Table 1.4b. This indicates therejection required at the input of LNA to prevent the reception of frequencies outside thereceive frequency band and limit the noise bandwidth of LNA. Rejection is also requiredover transmit band to prevent saturation of LNA by the high power transmit signals.

1.1.5.18 COMMANDABLE GAIN IN FGMThere should be provision for gain setting of each channel by ground command, by varyingattenuators in the transmission path. The minimum value of command-able gain shall be asper the specification (refer Table 1.4b).




There should be direct commanding of the targeted gain step and these commanding changesshall not introduce gain excursions outside the level range defined by the beginning and endstates.

1.1.5.19 ALC MODE SPECIFICATIONSThe specifications for minimum ALC mode dynamic range and output power controlattenuator (OPCA) shall be as per Table 1.4b. OPCA shall be capable of driving TWTA overa range varying from at least 12 dB back-off drive level of TWTA to 3 dB above thesaturating TWTA drive level. The ALC mode output power variation over the specified inputdynamic range shall be as specified in Table 1.4a.

1.1.5.20 FREQUENCY CONVERSIONThe receive frequencies shall be translated to transmit frequency band by conversionfrequencies as mentioned in Table 1.4b. The frequency conversion performance as well as thephase noise induced on single un-modulated carrier shall not exceed the values as given inTable 1.4b.

1.1.5.21 LINEARITY OF RECEIVE SECTIONWhen the spacecraft is illuminated with two equal amplitude carriers, the ratio of carrier toeach third order inter-modulation product in the receive section shall be as given in Table1.4b. The receive section is defined as the part of transponder between the receive antennaoutput and the HPA input.

1.1.5.22 IN-BAND FREQUENCY RESPONSE (SATURATION)The gain flatness of the transponder forward and return channels should not exceed the valuesgiven in Table 1.4b. This specification applies when transponder is in saturation.

1.1.5.23 IN-BAND FREQUENCY RESPONSE (SMALL SIGNAL)The gain flatness of the transponder forward and return channels should not exceed the valuesgiven in Table 1.4b. This specification applies when transponder is in small signal (linear)operation.

1.1.5.24 INPUT GAIN SLOPE RESPONSEThe maximum gain slope within the usable bandwidth of any transmission channel shall notexceed the values given in Table 1.4b. Input gain slope is measured between the input to thetransponder and the input to the high power amplifier and includes the contribution of thereceive antenna.

1.1.5.25 TOTAL GAIN SLOPE RESPONSEThe maximum gain slope within the usable bandwidth of any transmission channel shall notexceed the values given in Table 1.4b. The total gain slope is the performance of thecomplete transponder channel including receive and transmit antennas.




1.1.5.26 IN-BAND GROUP DELAY (INPUT)The Group Delay specifications (refer Table 1.4b) shall be met over the usable bandwidth.All the group delay values are relative with respect to the value at the center of the channel.

Input group delay is measured from the transponder input to the input of the high poweramplifier inclusive of receive antenna.

1.1.5.27 IN-BAND GROUP DELAY (TOTAL)The Group Delay specifications (refer Table 1.4b) shall be met over the usable bandwidth.All the group delay values are relative with respect to the value at the center of the channel.The Total Group Delay shall be within the specified limits for any illumination flux densitiesfrom saturation to linear operation of the channels. Total group delay is measured through thecomplete transponder, inclusive of transmit and receive antennas.

1.1.5.28 OUT OF BAND RECEIVE NARROWBAND REJECTIONSThe Receive Out-of-band Rejection relative to both the upper and lower band edge of the Ka-band channels shall be as per the specifications given in Table 1.4b. This parameter shall bemeasured from the transponder input to the input of the high power amplifier inclusive ofreceive antenna.

1.1.5.29 OUT OF BAND TRANSMIT NARROWBAND REJECTIONSThe Transmit Out-of-band Rejection relative to both the upper and lower band edge of theKa-band channels shall be as per the specifications given in Table 1.4b. This parameter shallbe measured from the HPA input to the transponder output including transmit antenna.

1.1.5.30 TOTAL CHANNEL PHASE-SHIFTThe total phase shift of a carrier transmitted within the usable channel bandwidth of anycommunications channel shall be as per the specification as given in Table 1.4b. This is thephase shift measured at the input power levels corresponding to specified IBO with respect tophase shift measured at 20 dB IBO. It shall be measured over the complete transponder chain.

1.1.5.31 AM/PM TRANSFER & CONVERSION COEFFICIENTThe transponder channel AM/PM Transfer and Conversion coefficients shall not exceed thelimits shown in Table 1.4b for any illumination level up to saturation. The transfer coefficientis measured with two carriers having an amplitude difference of up to 20 dB, with the largercarrier amplitude modulated to a depth of 1 dB peak to peak.




1.1.5.32 SPURIOUS OUTPUT

1. Spurious Within Transmit Band: The total RMS power, resulting from the sum ofall spurious signals (excluding the spurious signals resulting from uplink carrier,harmonics of the local oscillator frequency, and discrete spurious due to HPA powersupply), in the transmit frequency band, measured at the input to the transmit antennashall be below the noise power in a 4 KHz band.

The RMS power of the spurious signal from the mixer inter-modulation productsresulting from the one uplink carrier and any harmonic of the local oscillatorfrequency shall be as per the specification (refer Table 1.4b), measured at the input tothe transmit antenna.

The sum of all discrete spurious signals resulting from HPA power supply, measuredat the input to the transmit antenna, shall be as per the specification (refer Table 1.4b),for any drive level from 20 dB below saturation up to saturation.

2. Spurious Outside Transmit Band: The total RMS power of all spurious signalsoutside the transmit band and also output power due to harmonics of the normaloutput signals measured at input to the transmit antenna shall be as per thespecifications (refer Table 1.4b).

3. Spurious due to Receiver LO: The discrete signals within transmit band due to anyharmonics of Local oscillator frequency, when operated at maximum SFD, shall be asper the specifications (refer Table 1.4b).

1.1.5.33 OVERDRIVEAll the channels should be designed to withstand, for a minimum of two hours, single ormulti-carrier illumination flux densities up to 20 dB in excess of those required for thetransmission channel saturation at the lowest gain setting without subsequent degradation ofperformance. Other performance specifications may not meet in the affected channel duringthe condition of over drive. The qualification model should be subjected to at least 24 hrs ofoverdrive.

1.1.5.34 SWITCHING INTERRUPTSThere should be provision to recover service in case of HPA turning off due to any reason, tominimize total downtime.

The HPA should incorporate automatic restart function for handling micro-discharge events.All automatic restart events shall be reported through spacecraft telemetry. The HPA shall beswitched off automatically if a subsequent discharge or a sequence of discharges interruptsthe restart function, and it shall be restarted only by ground command.




1.1.5.35 TWTA CATHODE LIFETWTA cathode life should be ensured for at least 5 years in addition to the design life of thespacecraft.

1.1.5.36 UNIT INTERDEPENDENCEFailure of a single unit shall not affect any other unit. In case of HPA, failure of one of itsunits (i.e. EPC, TWT, LCAMP) shall only affect that particular HPA.

[Dual EPC configuration if to be used by contractor, shall not impact the reliability of theoverall system.]

1.1.5.37 HIGH POWER CONSIDERATIONSThe considerations related to the high RF power levels are Multipactor breakdown, PassiveInter-modulation Products (PIMP) and EMI/EMC.

a) MultipactionThe contractor should ensure the absence of any Multipaction related problems as theoutput section of transponder is subjected to high RF power levels. There should be nomultipactor breakdown in the subsystems when subjected to RF power which is at least 6dB higher than the maximum operating power level. This should be established throughanalysis and measurement.b) Passive Inter-modulation ProductsThe contractor should ensure the absence of any passive inter-modulation products in thehigh power components of transponder. This should be established by measurement. Theimpact of any passive inter-modulation product formed should be analyzed and confirmedto be negligible and acceptable.c) EMI/EMCThe contractor should ensure that the transponder and the individual subsystems shallmeet the EMI/EMC specifications as per MIL-STD-461E.

1.1.5.38 CESSATION OF EMISSIONSIt should be possible to turn each individual transmission channel ON or OFF by groundcommand.

1.1.5.39 PARAMETERS FOR THROUGHPUT ESTIMATIONContractor should calculate throughput using the following assumptions for the groundsegment. The calculations must consider the EIRP, G/T and aggregate C/I over 60% coveragearea as a reference. The link calculations must include adequate margins related to payloaddegradation, and modem implementation. Only the usable bandwidth and not the allocatedbandwidth should be used. The assumptions related to ground segment have been provided inTable 1.5a




Table 1.5a: Ground System Assumptions

Sr. No. Parameter Unit ValueHub Terminal

1 Antenna Diameter m 8.12 EIRP (incl. BPE and ULPC Margin) dBW 753 G/T (incl. BPE) dB/K 374 ULPC Margin dB 9

User Terminal5 Antenna Diameter m 0.86 HPA Sizing W 37 EIRP (incl. BPE and ULPC Margin) dBW 408 G/T (incl. BPE) dB/K 17.59 ULPC Margin dB 8

Data Rate Assumptions10 Forward Data Rate per Carrier Mbps 4011 Return Data Rate per User - As per Note

below*

* Note: The data rate per user in the Forward Link shall be 512 kbps. The number of users inForward link will be estimated accordingly. In the return link, data rate per carrier will beestimated based on the same number of users as the Forward link. The capacity must beprovided in clear sky as well as faded condition. The availability requirements and thecorresponding rain fade attenuation (in dB) are provided in Table 1.5b.

Table 1.5b: Availability and Rain Fade Assumptions

Sr. No. Parameter Availability AttenuationForward Link

1 Uplink 99.9% 30 dB2 Downlink 99.6% 7 dB

User Link3 Uplink 99.6% 15 dB4 Downlink 99.9% 14 dB

Appropriate fade mitigation techniques may be applied in the in the Forward link and thereturn link such as Site Diversity, Uplink Power Control (ULPC) and Adaptive Coding andModulations. For any assumption related to site diversity, appropriate reference of ITUstandard should be provided. The ULPC limits for user and hub terminals are provided inTable 1.5a.




The spectral efficiency and Es/N0 values for different mod-cods for link calculation are to betaken as per DVB-S2 standard in the Forward Link and the DVB-RCS2 (NG) standard in thereturn link. Appropriate margins for payload degradation, modem implementation and ACMloop adaptation should be considered.

Contractor may also provide alternative techniques for computing capacity in additionto that provided above.

1.1.6 KA-BAND BEACON SPECIFICATIONSTable 1.6 gives the specifications for Ka-band beacon transmitter.

Table 1.6: Ka-Band Beacon Performance Specifications

S. N. Performance Parameter Unit Specifications

1 Transmit Frequency GHz TBD & TBD2 Effective Isotropic Radiated Power (EIRP) dBW >143 Antenna Coverage - Service Area4 Polarization sense - TBD5 Cross Polarization Discrimination (XPD) dB >30

6

Frequency Stabilitya) Over a dayb) Over life time ppm

< ± 1< ± 10

c) Phase Noise : SSB level at frequencyoffset- 10 Hz- 100 Hz- 1.0 KHz- 10.0 KHz

dBc/Hz< -25< -55< -80< -90

7EIRP Stabilitya) Over any operating dayb) Over operating life

dB pp < 1< 2

The description of specifications is given below:

1.1.6.1 TRANSMIT FREQUENCYThe transmit frequencies of the Ka-band downlink beacons shall be as per Table 1.6.

1.1.6.2 EIRPThe effective isotropic radiated power (EIRP) specification of beacons shall be as per Table1.6.




1.1.6.3 COVERAGEThe coverage of beacons shall be service area so that it can cater to all the user and hubbeams.

1.1.6.4 POLARIZATIONThe polarization sense of the beacons shall be as per Table 1.6.

1.1.6.5 CROSS POLARIZATION DISCRIMINATION (XPD)The XPD of the beacon should be better than the value specified in Table 1.6 over specifiedcoverage.

1.1.6.6 FREQUENCY STABILITYThe frequency stability over the operating temperature and over entire life as well as thephase noise of transmitter shall be as per the specifications (Table 1.6).

1.1.6.7 EIRP STABILITYThe EIRP in the direction of any location within the coverage shall not vary by more than thepeak-to-peak values shown in Table 1.6 over any operating day and over lifetime of thesatellite. This specification includes antenna effects also. At the end of life, the specifiedEIRP stability shall include all degradation factors.

1.1.7 SPECIFICATIONS OF IOT-HORNTable 1.7 gives the specifications for Ka-band In Orbit Test (IOT) horn.

Table 1.7: Ka-Band IOT Horn Specifications

S. N. Performance Parameter UnitSpecifications

Transmit Receive1 Frequency Range of Operation GHz TBD2 Antenna Gain dBi > 22 > 22

3 Coverage - Service Area (ReferSection 1.1.2)

4 Polarization sense - TBD5 Cross Polarization Discrimination (XPD) dB >30

1.1.8 OPTIONAL CONFIGURATION (INDIA BEAMS + STEERABLE BEAMS)As an optional requirement, the contractor shall propose a scheme with four additionalindependently steerable beams. The steerable beams shall be used outside the specifiedcoverage to provide additional ad-hoc capacity.

The spectrum for each steerable beam will be 125 MHz. The hub spectrum required for thesebeams (500 MHz) shall be derived by reducing the number of user spot beams over specified




coverage from 88 to 86. The user spectrum for the steerable beams shall be same as the userbeams. Contractor may suggest strategies for payload configuration in congruence with this.

The payload configuration, and beam layout shall be finalized at the time of contract. Theuser beams to be switched off and the hubs catering to steerable beams shall be finalized atthe time of PDR.

1.1.8.1 PAYLOAD PERFORMANCE SPECIFICATIONSThe performance specifications applicable only for the steerable beams are given in Table1.8. The remaining specifications (including transponder performance) shall be same to theperformance specifications of the main payload (Refer Section 1.1.5).

Table 1.8: Payload Performance Specifications

S. N. Performance parameters Units Specifications

1. Beam size deg < 1.5

2. Beam steerability* - Over the CoverageRegion

3. Steering accuracy deg < 0.1

4. EIRP(single carrier saturation)over 100% beam area

dBW > 56

5.G/TOver 100% coverage area

dB/K > 8

6. Bandwidth per beam MHz 125$

7. Polarization sense (transmit & receive) - TBD

8. Cross Polarization Discrimination (XPD) dB >30

* The design & implementation of scanning system shall comply to steering of the beamsacross entire visible Earth Disc from the GEO slot.$ The usable bandwidth shall be at least 90% of this bandwidth.




2 TELEMETRY, COMMAND AND RANGING (TCR)SYSTEM

2.1 GENERAL REQUIREMENTSThe telemetry, command and ranging (hereinafter referred to as "TCR") subsystems shallprovide functions for controlling, monitoring and ranging the satellite. The TCR system shallbe capable of being operated any time during system level tests, pre-launch, launch phase andover the design Life, so that it meets all performance requirements.

A high gain antenna (either communication antenna or independent high gain antenna) shallbe used for “on orbit” normal operation in the geostationary orbit and when the requiredpointing accuracy of high gain antenna is not established (e.g. satellite attitude loss), wideangle beam antennas (Omni antenna system) shall be used. The switch over shall beaccomplished automatically without ground intervention.

The TCR system shall be CCSDS standard compatible.

The Telemetry (TM) and Command (TC) base-band systems shall have hot redundancy.

The TCR subsystems and the communications subsystem shall be capable of being operatedconcurrently without any interference.

The TCR signals shall respect the ITU regulations.

TCR high power amplifiers for transmission through wide angle beam antennas (omniantenna system) shall be independent from payload high power amplifiers.Any single pointfailure shall not affect the performance of the TCR system.

Compatibility shall be ensured and demonstrated between spacecraft telemetry, commandand ranging characteristics with the transfer orbit and on orbit ground control network. Acompatibility model (Suitcase model) identical to spacecraft TCR shall be used for thedemonstration.

The following analysis / document shall be presented for various reviews for acceptance.

1. Link budget analysisa. Separate tables for “Transfer orbit” operation, Nominal “on orbit” operation and

“on orbit” emergency operation for worst case and nominal case.b. A minimum margin of 3 dB shall be demonstrated under worst case (using

adverse value for every parameter).




c. Overall atmospheric attenuation (including attenuation due to rain) correspondingto 99% availability (within Indian region for on orbit operation and worst climatezone for Transfer orbit operation) shall be considered.

d. All assumptions shall be indicatede. The margins shall be demonstrated for a Command link BER of 10-5 and

Telemetry link BER of 10-6.f. For near omni coverage (transfer orbit link and on orbit emergency link), the

minimum antenna gain to be considered shall be for 99% coverage.g. An implementation loss of 2dB shall be considered for the link analysis.

2. TCR system design description

3. TCR RF compatibility analysis

4. TCR RF validation plan (equipment level and spacecraft level)

5. TCR Ground station Spacecraft interface (GSID)

6. TCR design verification matrix (DVM)

7. TCR Configuration Identification document list (CIDL)

8. GTD analysis

2.2 COMMAND SUBSYSTEM2.2.1 GENERAL REQUIREMENTSThe command subsystem shall be capable of being operated any time during spacecraft leveltest, pre-launch, launch phase and over the design Life, so that it meets all performancerequirements.

It shall not be possible to disconnect power or turn off any essential unit in the commandsubsystem while power is applied to the Spacecraft power bus. No bus transient, includingfuse burns, shall result in a non-recoverable change in the command subsystem operatingstatus or degrade the performance of the command subsystem to the point it becomesinoperative. Any expected worst-case bus voltage or environmental condition shall not resultin spurious command execution.

Measures shall be taken so that interferences from the communications subsystem or otherbus subsystems are not decoded and are not executed as command signals.




The design shall allow commanding via the omni antenna system at any time during thedesign Life, with no interference to and imposing no restrictions on the communicationspayload operations. The design shall be such that there shall be no impairment or restrictionwith emergency command uplink through omni antenna irrespective of state of any switchesin the command subsystem (including failed / switch does not contact either pole).

No switches of any kind shall be used in the omni path of the command subsystem fromantenna to the command decoder. There shall be no switches between command receiversand command decoders.

2.2.2 CONFIGURATIONThe command subsystem shall have more than or equal to 100% hot redundancy.

The command subsystem shall consist of at least two fully redundant paths, from the RFreception in the Spacecraft until the input of the commanded unit for the omni (wide anglebeam antenna) path.

The command subsystem shall consist of at least two fully redundant paths, from the RFreception in the Spacecraft until the input of the commanded unit, excluding only theantennas and passive microwave hardware for high gain antenna path.

There shall be minimum two command receivers (hot redundant) and these receivers shall becontinuously capable of receiving commands at any time.

There shall be two identical (hot redundant) command base-band systems and these systemsshall be continuously capable of receiving commands at any time.

It shall be possible to execute any command using either of the base-band systems. Loss ofone section shall not have any effect or degradation on other systems.

The command subsystem shall not require changing the command frequency whentransitioning between on-station command paths and emergency command paths (omniantenna system).

For on-orbit operations there shall be at least two antenna paths available at all times forcommanding.

The redundant command receivers and the redundant command decoders shall be cross-strapped.

In case of the satellite attitude loss, the command subsystem shall be switched to the wideangle beam antenna automatically.




2.2.3 FREQUENCY OF OPERATIONThe Spacecraft shall have two different command frequencies. The center frequency ofcommand signals shall be TBD GHz (in Ku Band). Each command receiver shall be operablewith unique single frequency.

TCR command subsystem shall be designed under the condition that the co-polar commandfrequencies are spaced by TBD MHz (minimum) with respect to each center frequency on thesame orbital slot spacecraft, using proposed command modulation parameters. TBD(frequency separation) shall be proposed by contractor.

Customer shall be allowed to modify the frequencies and polarizations up to Spacecraft PDR.

2.2.4 FREQUENCY STABILITYThe center frequency of command receiver shall be within TBD parts per million of nominalfrequency. TBD (frequency stability) shall be proposed by contractor.

2.2.5 ANTENNA COVERAGEThe antenna coverage shall be as follows:

1. If communication antenna is used for “on orbit” operations, then, with the pointingaccuracy of communications antenna satisfied on geostationary orbit, the coverage shallbe (mainland) India and it shall be provided by the contractor.If independent high gain antenna is used for “on orbit” operations, then, with the pointingaccuracy of spacecraft platform satisfied, the coverage shall be (mainland) India and itshall be provided by the contractor.

2. With the wide angle beam antenna:The antenna coverage shall be 4-pi streadian (near Omni coverage) and the Contractorshall demonstrate, by analysis, the compatibility of the antenna coverage with missionrequirements prior to achieving geostationary orbit and for on-station emergencyconditions. This analysis shall consider blockage of the antenna pattern by the spacecraftbody as well as multipath effects, if any.

2.2.6 ANTENNA POLARIZATIONThe command polarization shall be circular (TBD) for LEOP and “on orbit” emergencyoperation. TBD will be finalized at the time of PDR.

The command polarization shall be circular/linear (TBD) for “on orbit” operations. TBD willbe finalized at the time of PDR.




2.2.7 ANTENNA AXIAL RATIOThe wide beam angle antenna (omni antenna) axial ratio shall be less than 5 dB. The highgain antenna shall have axial ratio/XPD such that the polarisation discrimination is betterthan 30 dB.

2.2.8 INPUT POWER FLUX DENSITYThe command subsystem shall meet the required performance when command transmissionsfrom the ground illuminate the satellite at the following flux levels

1. With the pointing accuracy of communications antenna / independent high gainantenna satisfied on geostationary orbit from -103 dBW/m2 to -90 dBW/m2

With the wide angle beam antenna the flux level is -80 dBW/m²

2. Illumination at the command frequency with power flux density up to -60dBW/m²shall neither degrade the performance of the satellite nor shorten its life.

2.2.9 MODULATIONCommand modulation shall be FM / PSK-PCM(Contractor shall propose the FM deviation, sub-carrier and data rate (500 bps or higher))

2.2.10 BASEBAND REQUIREMENTThe Tele-command system shall be complaint with ESA packet Tele-command standard. Fullcompliance with the standard is required with respect to

a) Physical layerb) Coding layerc) Transfer layerd) Segmentation layer

The on board reception, processing and distribution of commands shall ensure that norestriction arise when the ground transmits commands of any types at the highest possiblerate.

The Bit Error Rate shall be better than 10-5 at the output of the Demodulator.

All commands except the commands used internally shall be verified by telemetry.Verification by means of telemetry shall be provided for all commands whether they are sentfrom ground or stored on board and released at later time, or generated autonomously onboard.




For each Tele-command following telemetry shall be provided. Tele-command data, TCreception / validity status whether Tele-command is successfully received and wasconsidered to be valid and auxiliary information sufficient to determine unambiguously thecause of Tele-command rejection with monitoring of TC command counter, aux register,execution register etc.

No Tele-command packet segmentation shall be implemented.

On board command routing shall be made only by MAPs, application IDs or by anyadditional word up to 8 bit.

The base-band system shall provide ON/OFF, DATA commands and the ON/OFF commandsshall not be used for any value setting.

The Tele-command base-band system shall provide on board autonomy function to maximumextent possible so that operator intervention in day to day operation of spacecraft isminimized. Command system shall provide all on board auto commanding functions (that isautonomy of commanding on board the spacecraft) as per customer requirement. The systemshall provide autonomy functions like time tag commanding, auto temperature control; eventbased commanding, configurable command block execution and battery management etc. Allon board spacecraft auto functions shall be overridden by ground command. All softwarefunctions shall have enable / disable function by ground command.

Any worst case bus voltage or environmental condition shall not result in any spuriouscommand execution.

Command format shall be capable of transmitting single command at time or block ofcommands in continuous mode.

Each command decoders shall have unambiguous unique addresses.

Commands that are irreversible which effect the safety of the satellite or its operational lifetime or those which may cause the loss of communication shall be clearly identified anddesignated as hazardous commands. Such commands shall require at least two consecutiveand independent commands to be uplinked from ground as TC. For example, “ARM ANDFIRE”. There shall not be any time limit between these commands issued from ground.

2.2.11 COMMAND ENCRYPTION (OPTIONAL)As a command channel safety, the command system shall be equipped with capability tooperate in secured (encrypted) mode only or encryption and authentication mode.




The command system shall incorporate an automatic function to switch to unencrypted(clear) mode for loss of earth recovery.

User programmable timer shall be provided which will return to clear mode in the event nocommand has been received with in predefined period from ground.

2.3 TELEMETRY SUBSYSTEM2.3.1 GENERAL REQUIREMENTSThe Telemetry subsystem shall be capable of being operated any time during spacecraft leveltest, pre-launch, launch phase and over the design Life, so that it meets all performancerequirements.

Each of the telemetry active subsystems (transmitters, high power amplifiers) shall beoperable independently in “ON” or “OFF” condition by ground command without anyrestriction.

The telemetry transmitters and high power amplifiers shall remain “ON”, following recoveryfrom any bus transient including a short circuit, which reduces the bus to zero volts for anylength of time.

The two redundant telemetry transmitters / high power amplifiers shall have the capability tooperate simultaneously, transmitting the information from the selected base band sections.

Loss of any of the building blocks (transmitters, high power amplifiers, passive microwavecomponents, RF switches and antenna) of the telemetry subsystem, in part or in whole, shallin no way affect the operation of the remaining section.

There shall be two identical (hot-redundancy) Telemetry (TM) base-band systems. Anysingle point failure shall not result in loss of any spacecraft data, that is, with one unit it shallbe possible to operate the full spacecraft throughout its designed life. Any single point failureshall not degrade the performance of other functional unit

2.3.2 CONFIGURATIONThe Telemetry subsystem shall have more than or equal to 100% redundancy.The Telemetry subsystem shall consist of at least two fully redundant paths, from thetransmitters to RF transmission antennas in the Spacecraft for the omni (wide angle beamantenna) path.

The Telemetry subsystem shall consist of at least two fully redundant paths, from thetransmitters to RF transmission in the Spacecraft, excluding only the antennas and passivemicrowave hardware for high gain antenna path.




There shall be minimum two Telemetry transmitters and these transmitters shall be crossstrapped to Telemetry encoders.

High power amplifiers for wide angle beam antenna shall be independent from payload highpower amplifiers.

The TM base-band systems shall have “ON/OFF” provision and it shall be possible to switch“ON/OFF” each of the TM base-band system independently.

The Telemetry format shall be CCSDS compatible.

The system format shall have programmable TM feature to be defined by user.TM system shall have provision of convolution coding as channel coding. The system shallhave include /exclude of coding by ground command for the down link.

2.3.3 FREQUENCY OF OPERATIONThe Spacecraft shall have two different Telemetry frequencies. The center frequency ofTelemetry signals shall be TBD GHz (in Ku Band). Each Transmitter shall have thecapability to operate in either of the frequencies. The frequency selection is by command.

TCR Telemetry subsystem shall be designed under the condition that the co-polar Telemetryfrequencies are spaced by TBD MHz (minimum) with respect to each center frequency on thesame orbital slot spacecraft, using proposed modulation parameters and taking into accountof the interference from another satellite's telemetry on the same orbit. TBD (frequencyseparation) shall be proposed by contractor.

Customer shall be allowed to modify the frequencies and polarizations up to Spacecraft PDR.

2.3.4 FREQUENCY STABILITYThe stability and offset of a telemetry carrier frequency shall be less than ± 1 part per millionduring any 24 hour period, under all conditions including eclipse.

The long-term frequency stability of telemetry transmitters, with all errors included shall bewithin ± 50 KHz of nominal frequency for all operating conditions of the satellite in orbit.The short-term phase stability shall be such that the phase noise of the un-modulated carrier(considering all the sources of spurious and noise within the spacecraft) shall not exceed5 degrees (RMS) when integrated between 100 Hz to 1000 Hz.




2.3.5 ANTENNA COVERAGEThe antenna coverage shall be as follows

1. If communication antenna is used for “on orbit” operations, then, with the pointingaccuracy of communications antenna satisfied on geostationary orbit, the coverageshall be (mainland) India and it shall be provided by the contractor.If independent high gain antenna is used for “on orbit” operations, then, with thepointing accuracy of spacecraft platform satisfied, the coverage shall be (mainland)India and it shall be provided by the contractor.

2. With the wide angle beam antenna, the antenna coverage shall be 4-pi streadian (nearomni coverage) and the Contractor shall demonstrate, by analysis, the compatibility ofthe antenna coverage with mission requirements prior to achieving geostationary orbitand for on-station emergency conditions. This analysis shall consider blockage of theantenna pattern by the spacecraft body as well as multipath effects, if any.

2.3.6 ANTENNA POLARIZATIONThe Telemetry polarization shall be circular (TBD) for LEOP and “on orbit” emergencyoperation. TBD will be finalized at the time of PDR.

The Telemetry polarization shall be circular/linear (TBD) for “on orbit” operations. TBD willbe finalized at the time of PDR.

2.3.7 ANTENNA AXIAL RATIOThe wide beam angle antenna (Omni antenna) axial ratio shall be less than 5 dB. The highgain antenna shall have axial ratio/XPD such that the polarisation discrimination is betterthan 30 dB.

2.3.8 EFFECTIVE ISOTROPIC RADIATED POWER (EIRP)The telemetry subsystem shall provide the following downlink EIRP:

1. With the pointing accuracy of communications antenna / High gain antenna satisfiedon geostationary orbit:

Telemetry EIRP shall be greater than or equal to + 15 dBW.

2. With the wide angle beam antenna:Greater than or equal to 0 dBW.

2.3.9 EIRP STABILITYIn geostationary orbit, the EIRP shall be stable within a 1 dB over any one day and 2 dB overthe Spacecraft Design Life.




2.3.10 MODULATIONThe primary transmission modulation for telemetry data shall be PCM-PSK/PM. Themodulation index shall be 1.0 0.10 radians. (Contractor shall propose the sub-carrierfrequency and data rate (1000 bps or more before coding))

2.3.11 BASE-BAND REQUIREMENTSThe status of any on board parameter of a sub system or its function shall be unambiguouslydetermined in the ground under all circumstances positively. The TM parameter shall monitorthe actual positional status of each relay/switch not its power status.

General categories of parameter to be instrumented and monitored via satellite telemetry shallinclude the capability to monitor the Analog, Digital, passive RF switch, Relay status,temp/pressure sensor, monitoring but not limited to these only.

TM system shall monitor the contents of memories or registers that determine the operationalstatus.

TM system shall monitor the data defining the causes of any automatic on board switchingand all significant information pertaining to the on board autonomy.

TM system shall monitor on/off status of all units, monitoring of load currents, temperatureinformation of all equipments whose temp is controlled automatically and status ofcommunication (P/L) unit consisting of RF switches, channel amplifiers, equipment status,RF drive level, TWTA helix current, anode voltage, antenna pointing and rotation forsteerable beams TWTA automatic restart occurrence status.

The design of analog TM monitoring shall be such that full scale range covers the fullperformance range of the parameter being measured, whilst ensuring that resolution isadequate for monitoring & evaluation purposes. Thos shall be ensured during all phases ofmission operations.

2.4 RANGING SUBSYSTEM2.4.1 GENERAL REQUIREMENTSA redundant ranging capability shall be provided with reception of ranging tones modulatedon the command carrier and re-transmission through Telemetry carrier.The ranging subsystem shall have more than or equal to 100% redundancy except forantenna.

The frequency of operation, antenna coverage and polarisation, the input flux densities andEIRP are same as indicated under Command subsystem and Telemetry subsystem sections




It shall be able to perform ranging without interrupting the transmission of telemetry signal(simultaneous Telemetry and ranging on a single downlink carrier).Simultaneous commanding and ranging shall not be required.

By ground command, it shall be possible to connect the ranging output of either of theCommand receivers to ranging input of either of the Telemetry transmitters (cross strapping).It shall be possible to “enable / disable” ranging functionality by ground command.The design of the on board Transponder (Receiver / Transmitter combination) shall becompatible with the ranging standard based on tones.

2.4.2 RANGE TONE FREQUENCIES AND MODULATIONThe ranging channel shall be compatible with tones selectable between 4 KHz to 100 KHz.

Modulation shall be FM for uplink and PM for downlink.

For ambiguity resolution, any delay difference between any minor tone and the major tonereferenced to the minor tone phase shall be less than ± 3°.

2.4.3 RANGE CALIBRATIONThe Spacecraft contribution to the TCR transponder ranging accuracy shall be less than 10meters (1 sigma) in the overall range measurement error budget.

The ranging tone stability shall be constant with in ± 80 ns over the full range of Doppler,input power flux density range, temperature range, spacecraft bus voltage and life.

Using the Telemetry values for input RF level reaching the onboard TCR transponder,temperature and predicted Doppler, it shall be possible to predict the transponder delay within ± 10 ns.

The Contractor shall provide a detailed ranging budget including all error sources for bothgeostationary and transfer orbit phases of the mission.

2.4.4 MODULATIONThe modulation for ranging shall be

Uplink : FMDownlink: PM

(Contractor shall propose the uplink FM deviation and downlink modulation index)

2.4.5 RANGING PERFORMANCE REQUIREMENTThe following shall be considered for the ranging operations.

A minimum major tone to noise density ratio is 27 dB-Hz




Implementation loss of 2 dB Minimum downlink modulation index of 0.45 radians for ranging tones TM sub-carrier + 2 tons of Ranging, the net index shall not exceed TBD (shall be

indicated by contractor so as to meet the link budget margins)

2.5 GROUND STATION COMPATIBILITY UNIT (SUITCASE MODEL)A TCR Ground station compatibility unit shall be provided to verify the compatibility of thesatellite with the ground stations used in the various phases of the mission. The Compatibilityunit shall be representative of all onboard TCR functions. The interface between theCompatibility unit and the G/S shall be through RF cables, variable attenuators and one mainAC input connection.

The TCR Compatibility unit RF frequencies shall be that of the spacecraft TCR frequenciesand if there are deviations from this, the customer has to be informed for acceptance of thedeviation.

The TCR Compatibility unit shall have independent and standalone functionality test facilityand procedure to verify the correct functioning of the unit on site.

The TCR unit shall be built with Engineering model standard equipments / modules.

The TCR Compatibility unit shall be delivered in a suitable transport container for ease oftransportation to ground stations.

The TCR unit shall be installed and made operational by contractor at customer’s premises.The contractor shall also provide adequate training to customer manpower for handling andoperation of the suitcase model

All functions shall be locally commandable and it shall be possible to monitor / verify allfunctionalities locally.

3 ELECTRICAL POWER SYSTEM3.1 GENERAL REQUIREMENTSThe electrical power subsystem of the spacecraft shall consist of all hardware and softwarerequired to generate, store, condition and distribute power to all loads in the spacecraftthroughout the mission.

The electrical power system shall support the specified payload and worst-case mainframesystems throughout the mission from launch up to end of operational life withoutinterruption.




The power subsystem shall perform its functions in the presence of all environments thespacecraft actually encounters, whether specified or not.

The electrical power system shall not impose any constraint on any of the mission operations.The primary source of power generation shall be photovoltaic solar arrays without anyconcentrators.

Energy storage shall be accomplished by Li-Ion batteries.

The pyro powering circuits shall be part of the power system and shall be operable under anycondition.

The power system shall provide the necessary provision for charging the batteriessimultaneously, at various rates from C/10 to C/200 with redundancy.

The power system shall be completely autonomous except for special operations like batterybalancing, battery cell bypassing etc.

No single fault either within the subsystem or elsewhere shall affect the nominal performanceof the subsystem or the spacecraft.

All automatic protection functions shall have manual override provision through groundtelecommand.

3.2 SOLAR ARRAYThe electrical power subsystem of the spacecraft shall use a photovoltaic solar array forpower generation.

The solar array shall consist of flight proven photovoltaic solar cells, protected with coverglass, laid down on composite substrates.

All components of the solar array and the processes shall be flight proven. Contractor shallconfirm that there are no in-orbit or in-testing failures or unexpected performance in solararrays of the same / similar design.

Use of concentrators of any design is prohibited.

No single point failure shall result in more than 5% reduction in power generation capacity ofthe array.

Allowable maximum inter string cell to cell voltages on the array is 30V.




Solar array shall be designed to survive the electrostatic charging environment encounteredduring the mission.

The solar array shall be designed to minimize charge build-up within the solar array andbetween the solar array and the various exposed parts of the spacecraft.The solar array designshall eliminate large structural currents between the solar array structure and SGRP.

Solar array design shall also minimize primary discharges between elements of a single solarcell assembly and eliminate secondary discharges between adjacent solar cells or solar cellsand the solar array structure.

All metallic parts of the solar panels, yokes and mechanism elements shall be grounded.There shall be no floating conductor part on the solar array.Any open circuit failure in thesolar array strings shall not lead to arcing/ discharges.

Contractor shall demonstrate compliance to the above requirements through test and analysis.Typical electron spectrum to be considered for test is shown in Table 3.1.

Solar cells, end terminations and solar array harness shall be protected with double insulationand this shall be demonstrated by test.Individual sub-strings of all solar array strings shall beisolated from other sub-strings of the same string by suitable isolation diodes.

All material and processes used in realizing the flight solar array including solar cells,substrate, harness, connectors, diodes and the laydown process should have been qualifiedthrough a qualification level thermal life test with 50% margin demonstrated for the designlife. If additional temperature cycling due to travelling shadow is predicted for anycomponent of the solar array, the thermal life test shall include this also.

Solar array design shall be magnetically clean. The residual magnetic moment of the solararray shall be calculated and accounted in the estimation of the platform pointing accuracy.

The trapped radiation fluence considered for the solar array design shall be based on the AP-8and AE-8 models. Contractor shall specify the variant used and the confidence level assumedfor the prediction.

The flare proton fluence considered for the solar array shall include as a minimum 6 ALflares of the Aug 1972 type.

Solar array performance and reliability shall not be adversely affected by the presence of theplumes from the chemical and electrical propulsion system. This shall be demonstrated byanalyses and if required by test.




Contractor shall provide as a part of the technical bid following information on the solararray:

a. Minimum inter-solar cell spacing on the array.b. Maximum voltage differential on the solar array with and without an open circuit

failure.c. Grounding scheme for the solar array.d. ESD mitigation measures to be implemented on the solar array and heritage of

such measures.e. Heritage of the solar array design proposed.f. Outline design of the solar array.

Table 3.1: Typical Electron Environment for Solar Array ESD Test

S.N. Energy, keV Density, pA/cm2

1. 20 10002. 50 3003. 100 1804. 150 605. 200 256. 250 127. 300 88. 350 5.59. 400 510. 450 411. 500 3

3.3 BATTERIESThe electrical power subsystem of the spacecraft shall use Li-Ion batteries for energy storage.

Batteries shall support all the spacecraft loads during launch, all eclipses and support thepower bus during the entire mission. Batteries discharge into the bus shall be automatic andcontrolled by the loads only. Automatic discharge cut-off shall be provided to preventcomplete discharge of the batteries.

The battery/ batteries shall consist of flight proven Li-Ion battery cells assembled in p-sconfiguration. All components of the batteries and the processes shall be flight proven.Contractor shall confirm that there are no in-orbit or in-testing failures or unexpectedperformance in batteries of the same / similar design.




The battery design shall be based on a reference energy (e) guaranteed to be delivered by itunder worst case in orbit operational conditions including charge & discharge profiles andbattery temperature.

This reference energy (e) and not the nameplate capacity shall be used for the power budgetestimation and DOD calculation.

During the mission, battery shall be maintained at the lowest required state of charge and itsEOCV shall be raised only to compensate for the capacity fading due to life and cycling.

Batteries shall be protected against open or short circuit failure:

a. Batteries shall be equipped with necessary bypass/ isolation circuitry which can beactivated for one cell at a time through telecommand from ground.

b. In case 1P batteries are planned, the bypass function shall be autonomous to preventan open circuited battery.

c. Single cell failure shall have no impact on battery performance.d. The bypass circuitry shall be protected against inadvertent operation through enable

and arm relays in addition to the fire relays.

Batteries shall be provided with a charge balancing mechanism to equalize State of Charge(SOC) of the cells (to within 10mv). The impact of cell voltage dispersion on availablecapacity should be accounted in the battery sizing.

Battery thermal design shall have the goal of enhancing reliability and life of the batterywithout compromising on performance. Battery thermal design shall allow the temperature tobe maintained within a narrow range during charge, discharge, charge balancing and bypassoperations.

Battery charging scheme shall allow the battery SOC at the beginning of the next discharge tobe, as a minimum, 97% of the reference energy (e).

The recharge rate of the batteries shall be variable through a ground telecommand from nocharge to the recommended maximum charge current.

No battery discharge for thruster firing

The charger circuit shall have sufficient capacity to allow the battery charging to becompleted within 18 hrs after a discharge to 75% Depth of Discharge (DOD).




Initiation of charge, tapering and termination of charge shall be automatic after each eclipsedue to the earth’s shadow.

For the battery discharges during the non-eclipse seasons, due to events like moon eclipses,charge current tapering and termination during post event recharge shall be automatic.Allowable maximum DOD (Based on reference energy) for the batteries with one cell/ cellmodule bypassed per battery shall be 75%. This shall be the baseline for estimating therequired battery capacity in the power budget.

There shall be no limitation imposed on the launcher for charging batteries during launch padoperations.

Design of the spacecraft shall allow removal of the batteries for storage without affecting thestate of qualification of rest of the spacecraft.

The spacecraft design shall provide access to mate and demate the spacecraft harness to thebattery and allow a battery servicing console to be interfaced to the battery.

Spacecraft design shall also permit charging of the batteries irrespective of the on/off statusof the spacecraft.

It is envisaged that the flight batteries will be part of the spacecraft during thermo-vac anddynamic tests of the spacecraft.

Contractor shall provide as a part of the technical bid following information on the batteries:

a. EOCV considered at the BOLb. EOCV upper limit at EOLc. Number of and nameplate & reference energy of the proposed batteries.d. Heritage of the proposed batteriese. Outline design of the batteriesf. Allocation of storage time

3.4 POWER CONTROL & PROTECTIONSThe design of the power system shall ensure that the bus voltage as experienced by the loadsshall be within the allowable limits specified, during all phases of mission, including thetransient when a battery cell is open or short circuited.

The contractor shall specify the bus voltage range for the various subsystems during variousconditions of spacecraft like normal operation in sunlight & eclipse, post battery emergencyin sunlight & eclipse and post AOCS safe mode.




Any single failure in the spacecraft, including those in components, harness and connectorsshall not result in either open or short circuiting of the bus.

The bus shall meet all the spacecraft requirements with the entire payload in synchronizedTDMA at any frequency from 0.1 Hz to 20 khz. If the proposed design of the power buscannot support synchronized TDMA of the entire payload, contractor shall specify themaximum number of TWTAs that can be supported within the allowable bus transients.

All spacecraft equipment shall be powered through redundant source end fuses such that anyfault in the equipment or harness between the bus and equipment does not propagate.

The power system design shall accommodate, as a minimum 2 spare fuse networks of eachtype for use after a fuse failure during the course of spacecraft testing. The fuse modules shallbe located such that, it shall be possible to interchange a spare fuse for a failed fuse withoutaffecting the qualification status of the spacecraft.

Upstream of the source end fuses, all power harness between the sources, bus and protectionfuses, including the SADA shall be double insulated.

All critical signals of the power system shall be identified and protected with doubleinsulation.

Distribution of “OR”ed secondary supplies of the power system is prohibited.

All high power payload equipment shall have under voltage turn-offs. No other payload orplatform equipment shall turn-off due to under voltage before these equipments.

The power system shall protect itself, the power and energy sources and the spacecraft loadsfrom the following:

a. Battery over discharge both in terms of voltage and currentb. Battery over charge both in terms of voltage and currentc. Bus over voltaged. Bus under voltage

If more than one battery is used, load sharing among the batteries shall be ensured such thatthe batteries have nearly same DOD.

The power system shall be entirely autonomous during all phases of mission except forspecial activities like lunar shadow, battery charge balancing and battery cell bypassing.




The pyro (EED) powering scheme shall have as a minimum, three physical breaks. Thescheme and the elements used shall be flight proven and shall meet the safety requirement oflaunchers identified.The power system shall include current sensors to measure load current, battery charge &discharge currents and solar array currents.

Power subsystem shall have sufficient telemetry for monitoring its health during testing phaseas well as in orbit. These shall include the requirements of the energy sources. These shallinclude but not limited to the following:

a. Solar array string inclusion statusb. Solar array voltages and currentsc. Charger status and currentsd. Battery voltages, individual cell voltages, charge & discharge currentse. Solar array, battery and power electronics temperaturef. Status of heaters of battery and power electronics unitsg. Battery EOCV set and charging statush. Load statusi. Solar array sun sensors errors and sun presence (If applicable)j. SADA temperatures & position monitoring

Power system telemetry shall provide sufficient observability to identify thefailure/malfunction of any load of the spacecraft.

Use of processor/ software for bus regulation is prohibited.

In case of sunlit regulated bus, bus de-latch from battery shall be ensured to be automaticwithout any disturbance to the loads.

Contractor shall provide as a part of the technical bid following information on the powerelectronics:

a. Heritage of the proposed power electronics subsystem.b. Number of buses and bus specifications.c. Outline design

3.5 REDUNDANCYThe solar array shall be divided into strings and for the purpose of power budget, failure ofthe string with the highest capacity shall be assumed.




If solar array sections are used for charging batteries, it shall be possible to charge all thebatteries simultaneously within 18hrs after the deepest eclipse, when one highest ratedsection per battery has failed.

The solar array design shall assume as a minimum, a 1% loss in current and voltage due torandom failures.

Even with two cells/ cell modules failed in the battery, a minimum of 5% margin in voltageshall exist with respect to the minimum bus voltage specified. If battery discharge regulatorsare used, this margin shall exist with respect to the minimum allowable input voltage for theBDR.

3.6 POWER BUDGETPower budget shall be made using measured power consumption figures of flightrepresentative units. Power budget shall be made for each of the power buses for solstices,equinox, eclipse and transfer orbit phases.

In the case of new units, a 5% margin shall be added to the estimated power consumptionfigures.

Power budget shall consider failure of one solar array string of maximum capacity per solararray wing.

If charger strings are used, the power generated by the charger section shall entirely beaccounted for battery charging during the eclipse season.

The power budget shall show a minimum margin of 7.5% for the solar array powergeneration on any day of the last year of mission considering the above mentionedrequirements.

The battery/ batteries DOD for the worst case earth and moon eclipse shall not exceed 75%considering one cell/ cell module bypassed per battery.

Contractor shall submit a power budget as part of the technical bid.




4 THERMAL REQUIREMENTSThe thermal control subsystem of the Spacecraft shall be designed to maintain units,subsystems and the entire Spacecraft within the specified temperature limits required toassure that the performance specifications are met during all stages of development andtesting before launch, during launch and post-launch under all mission conditions throughoutthe Design Life, including condition variances due to operation of a variable number oftransponders and due to actual Spacecraft electric power load conditions.

4.1 THERMAL RANGES AND TEMPERATURESAll unit / equipment temperatures shall remain within allowable limits during all stages oftheir development, integration, ground testing and during launch; and the on-orbit missionphase of spacecraft until end of operational life ,considering worst-case variations in powerdissipations, environments, operating modes, and contamination and degradation.

The extreme temperatures of all unit / equipment shall be generated by thermal analyticmodels. These values must be found unit by unit and / or equipment by equipment includingstructural elements, under a worst-case combination of thermal conditions.

A temperature margin shall be applied over the predicted temperatures from analyticalmodels to account for modeling uncertainties in order to arrive at worst-case acceptance orworst case qualification temperatures of each unit / equipment including structural elements.It is applied by adding the temperature margin to the maximum extreme or worst casetemperature and subtracting them from the minimum extreme or worst case temperature toarrive at acceptance or qualification test temperatures for each unit / equipment includingstructural elements.

The temperature margin for acceptance test conditions of all Spacecraft equipments includingstructural elements shall not be less than 5° C with respect to worst-case temperaturepredictions. Thermal margin for equipment which is outside of the Spacecraft body envelope– including, but not limited to, antenna reflectors, sub-reflectors, TTC antennas, Sensors etc.,shall not be less than 10º C with respect to worst-case temperature predictions.

The temperature margin for qualification test conditions of all Spacecraft unit / equipmentincluding structural elements shall be not less than 10° C with respect to worst-casetemperature predictions.

Temperature margin for equipment which is outside of the Spacecraft body envelope –including but not limited to antenna reflectors, sub-reflectors, TTC antennas, Sensors etc.shall be not less than 15º C with respect to worst-case temperature predictions.




Survival / turn-on temperature is defined as the temperature in which the equipment need notoperate within specification, but must not experience any degradation below specificationlimits when returned to the operating range. Survival / turn-on margins shall be not less than15° C with respect to worst-case temperature predictions.

The contractor shall provide the thermal design margins for qualification, acceptance anddesign with adequate justification and supporting documents in the response to the RFP. Adetailed thermal budget shall be submitted by the contractor.

4.2 THERMAL DESIGNThermal control shall be achieved by passive means, together with the use of active elements.The thermal control subsystem shall be designed such that no constraints are imposed to thesatellite system and its unit / subsystem in any mode of operation at any time. All powerdissipations used for thermal control design shall have been verified by measurements onflight-type equipments. The performance of the Spacecraft thermal design shall bepredictable by thermal analysis and verifiable by ground test.

The use of gravity sensitive thermal control devices (e.g. constant conductance heat pipes)shall not impede or restrict the testing of the satellite according to the requirements set forthin development, qualification and acceptance requirements.

The solar absorptance value for the second surface mirrors used in any temperature predictionat end of design Life shall be about 0.25 or as per mutually agreed data. However, the solarabsorptance value for any heat rejecting surface used in any temperature prediction over thedesign life shall be fully justified/supported by a detailed contamination analysis, on-orbitdata, and appropriate test measurements. The contamination study shall also include anexamination of ways to reduce the degradation.

The contractor shall provide the solar absorptance value for the second surface mirrors to beused in temperature prediction at end of design life for thermal radiators in the response tothis RFP.

The values of thermo-physical properties used in temperature predictions shall be fullyjustified and appropriate tests shall be conducted to derive properties where significantuncertainty exists. These tests shall be conducted using identical hardware in the sameconfiguration that is intended for use on the Spacecraft.

The failure of any single heat pipes / loop heat pipes or their interface, if used, shall notpreclude the operation of any transponder, within its RF performance specification.Furthermore, no components shall be outside of their respective qualification temperatureranges as a result of a heat pipe or heat pipe interface failure. The contractor shall provide




proposed heat pipe heritage, technical capabilities, weight per unit length, acceptance /qualification plan, performance degradation aspects etc. in the response to this RFP.Deployable thermal radiators, if used, in the thermal design, the failure of any single heatpipe or loop heat pipe or its any thermal interface shall not preclude the operation of anytransponder, within its RF performance specification. Furthermore, no components shall beoutside of their respective qualification temperature ranges as a result of a heat pipe or loopheat pipe or it’s any thermal interface failure. The contractor shall provide proposeddeployable thermal radiator data related to the heritage, technical capabilities, weight,acceptance / qualification plan, performance degradation aspects etc. in the response to thisRFP.

All heaters for the thermal control of any equipment or structures or appendages shall befunctionally redundant with automatically temperature control functionality. In case offailure of one heater, the thermal control shall be capable to maintain all satellite units /equipments within their operating temperature ranges.

Heaters shall be provided with a ground commanded temperature control limit programmablecapability and a ground command override capability. There should be a provision fortemperature sensor programmability.

All heater circuits shall be designed to prevent the predicted Spacecraft temperatures withuncertainty margins applied from exceeding the hot acceptance temperatures or going belowthe cold acceptance temperature in the event of a single failure, either open circuit or shortcircuit, of any part of the heater circuit or associated control circuitry.

Heater protection shall be implemented through automated redundancy on the Spacecraftunless it can be demonstrated by analysis that the predicted Spacecraft temperatures willremain within the above requirement for at least 30 minutes in all cases, in which case groundcommanded intervention can be implemented.

Heater circuits shall be designed such that the loss of a single layer of electrical isolation atany point along the heater circuit wire or at any point in the heater element where the currentin the circuit will not exceed the guaranteed opening current of a fuse or similar protectiondevice, the predicted temperatures of the Spacecraft with uncertainty margins applied fromexceeding the hot acceptance test temperatures or going below the cold acceptancetemperature.

The design and layout of redundant heater elements has to assure electrical and physicalisolation from each other, such that damage effects (e.g. burnout) of one element do notdamage or deteriorate its redundant counterpart.




Contractor has to specify the heater power density being used on different surface (e.g.aluminum honeycomb, heat pipe etc). The contractor shall provide a list with the heater typesincluding attachment method on the Spacecraft and with the associated power density.

Where heaters are used to limit the expected worst-case cold temperatures, such heaters shallbe redundant and oversized by about 25%. Where this is done the expected worst-case heaterset point, including worst case overshoot as justified by analysis, shall be used for the coldtemperature extremes for both acceptance and qualification testing.

In-flight telemetry temperature measurements shall be provided in sufficient quantity andwith a sampling rate sufficient to allow verification of proper functioning of the satellite andtimely detection of anomalies.

In case of failure of one temperature sensor, the thermal control shall be capable to maintainall satellite units / equipments within their temperature ranges. The contractor shall provide atentative location list of temperature sensors in the response to this RFP.

The temperature measurement points shall coincide with nodes defined in the thermalmathematical model. All temperature sensors for the thermal control of any unit / equipmentor structures or appendages shall be functionally redundant. The adequacy shall be suitablyverified during thermal vacuum testing, if required.

A power requirement of thermal system is very crucial and shall be minimized during allphases of design life. The contractor shall indicate the expected thermal power requirementfor each major subsystem and the overall thermal power requirement at launch pad, transferorbit and various season at beginning of life and end of design life of in-orbit phase in theresponse to this RFP.

The weight of thermal system shall be restricted to lowest possible (about 6 % of the total dryweight of the spacecraft). The contractor shall indicate the expected weight of all thermalelements and the total weight of thermal system in the response to this RFP.

4.3 THERMAL MATHEMATICAL MODELS AND ASSOCIATEDDOCUMENTATION

Designs of the thermal control subsystem of the Spacecraft and its subsystems shall besupported by adequate thermal analysis during all stages of development and testing beforelaunch, during launch and post-launch under all mission conditions throughout the DesignLife, including condition variances due to operation of a variable number of transponders ,Spacecraft electric power load conditions, each season at both BOL and EOL and includediurnal effects, eclipse effect etc. as applicable.




The size of thermal mathematical models shall be adequate to have thermal analysis accuracyand contractor shall use reasonable best efforts to include as nodes in the thermal model asmany flight temperature sensors as practicable.

The thermal effects of individual unit (e.g. high power isolator) or transponder channelfailures shall also be considered. Special care must be taken with analyses of interfaces withcritical satellite elements (e.g. antenna reflectors, solar arrays).

Contractor shall provide complete thermal design analyses of the spacecraft, its subsystemsand units over all expected mission conditions, including condition variances due tooperation of a variable number of transponders and actual Spacecraft electrical loads.

Temperature predictions during all stages of development and testing before launch, duringlaunch and post-launch under all mission conditions throughout the Design Life, includingcondition variances due to operation of a variable number of transponders, Spacecraftelectric power load conditions, each season at both BOL and EOL and include diurnaleffects, eclipse effect etc. as applicable, shall be provided at an appropriate time.

The expected thermal gradients across all propellant tanks at both summer and wintersolstices and equinox; and expected thermal gradients along all propulsion lines; expectedtemperature gradients in antennae , solar arrays etc. shall be provided in the reviewdocuments of thermal control subsystem.

The analytical thermal model used to predict the extreme temperatures at the spacecraft levelperformance shall be validated during the integrated spacecraft level thermal tests oversufficient different test conditions.

Correlation of the test results to the thermal model predictions shall be within 3° C.Technical explanation shall be provided where any unit observed temperature data exceedsthe cold or hot extreme predicted by more than 5° C.

In the event that there is any resulting uncertainty greater than the unit acceptance testmargin, the unit acceptance and qualification margins shall be increased accordingly. Theincrease in unit qualification margin shall be accomplished by test of representative units atthe earliest but not later than pre-shipment review clearance for thermal control subsystem ofthe spacecraft.

Following Thermal analysis documents, but not limited to, are required to be delivered:a. Integrated Spacecraft thermal analysisb. Equipment panel and radiator panel level thermal analysis.




c. Propulsion subsystem (s) including engine / thruster and all other elementsthermal analysis during all phases of the mission

d. Battery modules with its radiator thermal analysis.e. Power Switching Regulator thermal analysis.f. Solar panel thermal analysis.g. Thermal Vacuum Test thermal analysish. Antenna reflector and support structure including deployment mechanism thermal

analysis.i. T&C antenna thermal analysis.j. Solar array drive thermal analysis.k. High power RF component thermal analysis.l. Overall high RF power handling output section thermal analysis.m. Earth Sensor thermal analysisn. Star Sensor thermal analysiso. Analysis for Plume heat flux from liquid apogee engine and other propulsion

thruster on to satellite surfaces and appendages like solar arrays, deployableantennae etc. and its thermal effect.

In addition to the above, contractor shall provide Spacecraft thermal model in the format andlevel of detail required by the Launch Services Contractor. This model will be used byLaunch Services Contractor in the overall Launch Phase Spacecraft Thermal Analysis.

4.4 THERMAL SUBSYSTEM DEVELOPMENT TESTINGQualification / heritage of each thermal control element / processes / design tools for themission life shall be demonstrated by suitable test or by providing necessary documents, ifalready qualified.

The thermal optical characteristics of all surface finishes shall be verified by measurement.

Following thermal interface characterization tests, but not limited to, shall be demonstrated:a. Heat pipe to heat pipe,b. Unit / equipment to heat pipe,c. Heat pipe to radiator,d. Heat pipe dry out flux,e. Various thermal joints of deployable thermal radiators, if applicable.f. Unit/equipment to deck

4.4.1 ANALYSIS / MEASUREMENTSensitivity and integrity of thermal interfaces when exposed to:

a. Vibration loads,b. Heat flux,c. Temperature extremes,




d. Temperature rates of change,e. Aging,f. Material and manufacturing process variation,g. Bent pipe / straight pipe interface test,h. Variations in internal heat pipe pressures (achieved at thermal extremes),i. Differences in coefficients of thermal expansion for dissimilar materials,j. Differences in expansion across interfaces resulting from temperature differences.

4.4.2 HEAT PIPE / LOOP HEAT PIPE DEVELOPMENT TESTS:a. Unit level: Thermal performance, pressure tests, thermal soak / thermal cycling,

structural, Non-condensable gas generation, etc.

b. Heat pipe / Loop heat pipe assembly: Thermal performance including panel-to-structural, dynamic vibration etc.

4.4.3 DEPLOYED THERMAL RADIATORa. Unit level: Thermal performance, pressure tests, structural, non-condensable gas

generation, etc.b. Heat pipe / Loop heat pipe assembly: Thermal performance including

panel-to- structural, dynamic vibration etc.

4.4.4 SPACECRAFT THERMAL BALANCE TEST:The validation of the Thermal Mathematical Model (TMM) shall be through the thermalbalance test. Test simulation plan shall be mutually agreed before conduct of the test.Contractor shall use the predictions for cases defined as thermal balance cases to perform thecorrelation of the Spacecraft thermal model.

The following stability criteria or as per mutually agreement shall be applied for identifiedunit / subsystem:

“Less than O.2°C / hour during 5 hours for steady state”;

The criteria for success of the validation of the Thermal Mathematical Model, as a result ofthe correlation with the thermal balance vacuum test measurements shall be:

a. An average temperature deviation less than 3°C for equipment units;b. A maximum deviation less than 5°C in every point.

Contractor shall also provide temperature predictions for each phase of Spacecraft ThermalVacuum Performance Test prior to the test.

4.5 THERMAL CONTROL MATERIAL AND PROCESSESThe complete list of Thermal Control Materials and thermal processes used shall be providedby the contractor along with their heritage and qualification status.




5 STRUCTURE SUBSYSTEM5.1 FUNCTIONAL REQUIREMENTS:The structure subsystem shall provide the mechanical support for the other subsystems in aconfiguration which meets all the requirements of the satellite and the subsystems e.g.thermal control, mass properties, alignment, launch vehicle interface/requirements, assembly,integration and test etc.,

Structure subsystem of spacecraft shall preferably consist of two modules viz., bus moduleand payload module. Design and realization of modules shall be in such a way thatintegration of respective subsystems with the modules, handling, transportation etc., can bedone independently as separate modules and shall have the interface compatibility for thefinal assembly of the spacecraft.

5.2 PERFORMANCE REQUIREMENTS:The spacecraft structure shall be capable of sustaining all direct and cumulative static anddynamic load combinations occurring during fabrication, testing, ground handling,transportation, launch, in-orbit maneuvers and exposure to in-orbit failures (includingreaquisition), without exceeding the limit of elastic deformations. The interfaces with thelaunch vehicle shall ensure a correct separation.

The structure shall maintain throughout the design life the dimensional stability andalignment relationships required to enable all functional operations of the satellite and itssubsystems.

The structure shall also provide protection to other subsystems against excessive loads duringtesting, ground handling, transportation, launch, on-orbit deployments and maneuvers.

5.3 DESIGN REQUIREMENTS:5.3.1 ENVIRONMENTAL LOADSThe spacecraft structural design shall be compatible with the launch, ground processing, testand on-orbit environments of the spacecraft. Maximum loads shall be computed as the worstcase loads coming from the appropriate combination of these different loading conditions.

The structure shall be designed for the worst case combined envelope of environmental loadconditions, fracture mechanics and safety requirements associated with the designated launchvehicles as specified in Schedule E.

Structural stiffness criteria shall include considerations of minimizing dynamic coupling withthe launch vehicles as well as interaction of appendages (deployable like antennae, solarpanels etc.) with the satellite attitude control system.




The launch vehicle agency shall conduct the coupled satellite/launch vehicle dynamicanalysis from which the final flight limit loads shall be determined.

Spacecraft shall comply with the sine, acoustic, random and shock loads specified in the usermanual of each compatible launch vehicle.

5.3.2 DESIGN CRITERIAStructural elements shall be designed for maximum loads, which shall be multiplied by thesafety factors as specified in Table 5.1.Design loads shall be calculated as followsDesign load = maximum load X safety factor

Table 5.1: Safety Factor

S.N. Structural elements Safety factor

1. For primary structure which has been proof tested 1.32. Primary structure 1.4

3. Equipment panels, secondary structure,appendages and equipment

1.5

4. All items subjected to repair under MRB control 1.5

The design loads shall be used to compute stresses. Computed stresses shall be compared toyield and ultimate material stresses and buckling loads or stresses of structural elements tocalculate margin of safeties.

Margin of safety = (Allowable stress/Design stress) - 1

The minimum required margin of safety based on failure mode is specified in Table 5.2.

Table 5.2: Margin of Safety

S.N. Failure mode Margin of safety

1. Metallic yield >0.102. Metallic ultimate >0.253. Composite ultimate

(First ply failure)>0.30

4. Local buckling >0.105. General buckling >2.00




For structures made from composites the margin requirements shall be met relative to anallowable stress or load taken as the minimum guaranteed material property data, obtainedfrom tests conducted on coupon samples of each actual lay-up used to build the flighthardware. An allowable value with suitable and appropriate failure criteria shall be used inthe design to estimate the margin of safeties. An allowable in this context is the value for themechanical property where at least 99% of the population of values is expected to fall with aconfidence level of 95%. Residual stresses in composite laminates, induced by cooling fromthe cure temperature shall be included in determining the ply stresses.

For composite and bonded structures, or any other structural elements affected by thermalcycling, moisture, cracking, ageing and radiation, the margin of safety shall be demonstratedafter their design strength and dimensional stability have been appropriately de-rated toaccount for these effects.

For structures made from low fracture toughness materials and for pressure vessels, thesatellite contractor shall implement a fracture control plan which shall establish formalacceptability criteria for flaws and fractures, based on flaw growth analyses, thresholdfracture toughness and appropriate inspection and test techniques to ensure compliance withlaunch vehicle contractor requirements.

Where higher margins of safety are required by launcher agencies or where local authoritiesimpose higher requirements during testing, transportation or preparation of the spacecraft,these higher requirements take precedence over the requirements in this section.

The spacecraft design shall meet the stiffness criteria consistent with the structural frequencyconstraints imposed by the launch vehicle agency.

Where heat pipes are used, the structure shall be designed such that the heat pipes are notneeded for carrying loads or required for meeting the launch vehicle agency imposed stiffnessrequirements.

All pressure vessels shall comply with the launch vehicle agency safety requirements.

The structural/mechanical subsystem design shall respect requirements for alignments andmechanical tolerances of all subsystems, including thermal distortion, hygro-elastic distortioncreep, residual yield and joint relaxation.

All bolts shall meet suitable aerospace standards, the use of machined bolts is strictlyforbidden.




6 ATTITUDE DETERMINATION AND CONTROLSYSTEM(ADCS)

6.1 DEFINITION OF SPACECRAFT AXESThe Spacecraft right hand orthogonal axes are defined as fixed in the Spacecraft, with theorigins at the nominal center of the Spacecraft to launch vehicle separation plane. When theSpacecraft is in the operational configuration in geosynchronous orbit, the Yaw axis positiveis in the orbital plane and directed toward the center of the earth; the Roll axis positive is inthe orbital plane in the direction of orbital motion; and the Pitch axis positive is normal to theorbital plane and directed towards the south.

6.2 ATTITUDE CONTROL SUBSYSTEM FUNCTIONALREQUIREMENTS

The attitude determination and control subsystem (ADCS) shall include all satellite hardwareand software required to determine and control the attitude and antenna pointing of thesatellite during all phases of the mission lifetime

The spacecraft shall have minimum two on board computers for the purpose of attitudedetermination and control. All components used to control the attitude of the spacecraft shallhave redundancy. Control wheels, if used, shall be redundant such that, in the event of failureof one wheel the other wheel(s) shall provide an equivalent function. Cross strapping betweenredundant sensors and actuators shall be provided. The redundant elements shall beelectrically and mechanically isolated.

It shall be possible to switch by ground command between all redundant ADCS equipment,without degrading the performance specified in this schedule during or following theswitching operation throughout the operational life. It shall be possible to command ON forall redundant ADCS equipments for health check while maintaining the primary chain incontrol, and without affecting the performance of the ADCS.

The ADCS shall have ground override capability to enable or disable every autonomousfunctions and logics onboard like firing of thrusters for attitude control for momentumunloading etc.

6.3 ATTITUDE DETERMINATIONThe spacecraft shall be capable of providing all necessary information to the ground controlstation for determining the inertial orientation of the spacecraft for all phases during thedesign life. It shall include the capability for calibration and correction of attitudemeasurements made by on-board sensors.




The attitude determination and control specifications shall be satisfied when the spacecraftis co-located with several other satellites, in particular considering that these other satellitescan orbit around the spacecraft and can be at any position in the orbital control window.

The ADCS shall be capable of accurately calibrating the gyro drift rates (better than0.025°/hr), and implement the resulting compensation, without any input related to thecurrent spacecraft attitude and ADCS sub-mode.

Sun and moon interference prediction on ground shall be with an interval of upto one month.The ADCS shall be capable to autonomously neutralize sun and moon interference effects byusing uplinked apriori predicted data.

The spacecraft shall be designed to autonomously estimate and maintain three-axis attitude aswell as estimate and compensate external disturbance torques.

6.4 ATTITUDE CONTROLThe spacecraft shall be capable of maintaining its attitude and antenna pointing within theaccuracy required to meet the specifications documented throughout the operational life. Theprimary antenna-pointing mode shall be by RF beacon tracking. Outside of momentumunloading maneuvers, station keeping maneuvers, and contingency operations, the ADCSshall be capable of maintaining autonomously the spacecraft pointing without firing thrusters.Three-axis stabilized spacecraft is highly desirable.

The attitude control torque capability shall be at least double the worst-case disturbancetorques and these shall be demonstrated taking into account all transient effects of thespacecraft dynamics, apogee engine fire, and control loops, including both step and harmonicperturbations over the control bandwidth.

The ADCS design shall include capability to recover from any flat spin or tumbling conditionthat could result from a temporary loss of attitude control in any of the mission phases.

In all control modes, including safe modes, status or data from all sensors used for on-board,closed loop attitude control shall be checked by the ADCS to detect failures, isolate andreconfigure (FDIR) in order to prevent the use of failed equipment.

6.4.1 STABILITYThe spacecraft ADCS shall be designed to guarantee adequate control loop stability marginfor all ADCS discretized control loops with the nominal and worst case configurations, andtaking into account all the discretization effects. The stability analysis shall also take intoaccount all the contributors that could play a significant role in the performance and/or thestability, including delays, non-linearity and flexible/sloshing modes.




The flexible/sloshing modes shall be rejected to the extent possible, in case it would not bepossible, then these modes should be necessarily controlled. The design/tuning of the controlloop used for thruster control shall minimize the fuel consumption while meeting all themission requirements, performance and stability requirements.

6.4.2 POINTING ACCURACIESThe spacecraft pointing shall be maintained within the limits set forth in Table 6.1 & 6.2 asgiven below. The contractor shall demonstrate that the pointing limits in Table 6.1 & 6.2 areconsistent with the allocations for the payload beam pointing requirement.

Note to contractor: Contractor may note that table 6.1 & 6.2 shows spacecraft level pointingrequirements; however payload beam pointing shall be met by appropriate methods.

Table 6.1: Pointing Requirements (Without Electric Propulsion System)

S.N. SpacecraftAxis

Normal mode, includingMomentum Unloading,degrees.

Station Keepingmode, degrees

1. Roll 0.10 0.15

2. Pitch 0.10 0.15

3. Yaw 0.20 0.20

Table 6.2: Pointing Requirements (With Electric Propulsion System)

S.N. SpacecraftAxis

Normal mode,includingMomentum Unloading,degrees.

Station Keeping mode,degrees

1. Roll 0.10 0.10

2. Pitch 0.10 0.10

3. Yaw 0.20 0.20




6.4.3 STATION KEEPING MANEUVERS:The ADCS shall be capable of maintaining Spacecraft pointing without constraints on stationKeeping maneuvers due to the position of the solar wings relative to the body, or due to theselected mode (primary or redundant thrusters for delta-V), or due to sensor selection or itsmode of operation.

If the station keeping maneuvers are being performed using low-thrust Electric Propulsionsystem, the ADCS shall maintain the Spacecraft pointing without the use of chemicalthrusters.

In case of the usage of Electric Propulsion System for Station Keeping Maneuvers, a BackupOption with Chemical Propulsion shall be provided

The satellite design shall allow inclination correction maneuvers to be performed in both theNorth and South Directions

6.4.4 IN-ORBIT ANTENNA PATTERN MEASUREMENT:The ADCS shall be designed to support the Antenna Pattern Measurement. The proposedstrategy shall allow a sufficient range of satellite attitude variation, and sufficient attitudedetermination or reconstitution accuracy to meet the requirements of Communication PayloadRequirements, in a propellant efficient manner.

The satellite attitude variation shall be pitch cuts at specific roll offsets (maximum ranges).The maximum rate of change of pitch during a cut shall be 0.02°/sec .The attitudedetermination or reconstitution accuracy shall be better than 0.03° and the relative accuracybetween the beginning and end of each pitch cut with an absolute accuracy of better than0.05° for all cuts.

6.4.5 CONTROL BIAS CAPABILITY:The satellite shall include means of biasing, by ground command, the zero-error settings ofthe attitude control subsystem for the purpose of accommodating the satellite at differentorbital locations, adjusting the antenna coverage or compensating for small orbit inclinations.As a minimum, it shall be possible to:

a) Bias the pitch attitude to within plus or minus 0.01 degree of any desired value up to± 6.0 degrees, whilst not degrading the specified pointing accuracy.

b) Bias the roll attitude to within plus or minus 0.01 degree of any desired value up to atleast ±2.0 degrees, whilst still meeting the specified pointing accuracy, whileminimizing any impact on the operational lifetime consistent with the ADCShardware.




The maximum safe roll and pitch bias shall be characterized as part of the flight equipmenttesting.

It shall be possible to set up a sinusoidal varying attitude bias in three axes with period equalto that of the orbit, in addition to a fixed component, without the need for repeated groundintervention. The sinusoidal attitude bias shall enable the compensation of antenna coverageand polarization correction during inclined orbits up to 3 degrees.

6.4.5.1 OPERATION IN INCLINED ORBIT AT THE EOLThe spacecraft shall be designed to autonomously maintain its antennae pointed to theirnominal position on earth when operating in inclined orbit up to a maximum of 3 degrees.

6.4.5.2 GROUND CONTROLThe satellite attitude and antenna pointing control shall be command-able from the ground asa backup in the event of multiple failures of reference sensors. It shall be possible to controlby ground command, all thruster firings, reaction or momentum wheel speeds, motion of slipring assemblies, magnetic or solar torque devices, and solar panel orientation.

6.4.5.3 UNIT AND REFERENCE SWITCHING:It shall be possible to switch by ground command between functionally redundant ADCSequipment, without generating attitude transients which cause the satellite to exceed themaximum allowable pointing errors. In addition, where normal operation of the subsystemrequires switching of control modes, control parameters, attitude references (e.g. for eclipseor to avoid sun/moon interference, etc), such switching shall not result in transients whichexceed the allowable pointing errors.

6.4.6 SAFE MODES AND RE-ACQUISITION:The ADCS design shall include safe modes and the appropriate safety features to ensure thesafety of the satellite at all times under failure conditions. In all operational modes deviationsfrom the desired attitude shall be minimized in case of a failure. When switching to redundantunits/modes, propagation of wrong data shall be avoided.

For the on orbit phase, the design shall ensure that any single sensor or actuator failure(excluding catastrophic scenarios, e.g. Seizure of wheel, gross propellant leak) shall not resultin loss of earth pointing and the deviation from nominal pointing requirements under suchcircumstances shall be of the order of less than 5 minutes duration.

A last resort level of safe mode may use the sun rather than the earth as pointing reference. Inthis case a recovery to normal earth pointing shall be as short as possible, typically 2 to 3hours.




At any point of time ADCS shall ensure that the spacecraft rates shall not exceed a prefixedvalue that maybe detrimental to the spacecraft subsystems including power generation andthermal management.

In safe mode the spacecraft shall not make use of any attitude determination and controlequipment (either sensors or actuators) that were in use before switching to safe mode sincethat could be considered as a possible cause of the switch itself.

The ADCS shall ensure that solar panels are sun tracking even in safe mode to ensureadequate power generation till recovery.

The ADCS shall ensure that no two automatic changeovers of the same subsystem take place.The second changeover should be after ground intervention/ground normalization only.Tele-command and telemetry shall be ensured via Omni antenna during safe mode.

6.4.7 PAYLOAD SHEDDING IN EMERGENCY:The safe modes of the Spacecraft shall be designed in such a way to minimize theinterruption of the communications traffic. For such to be accomplished, in all cases theautomated response to the fault, when such response includes shutting OFF the payload orpart of it, shall include the capability of disabling the payload shut down, while stillmaintaining the remaining fault correction actions. The fault response system shallincorporate alternatives to the payload shedding that allow faster recovery of the traffic. Theautomated payload shedding functions shall be presented at the design reviews and theirimplementation shall be subject to Customer’s approval.

TC & TM operations shall be guaranteed during large Spacecraft de-pointing, regardless ofthe payload reconfiguration.

6.4.8 MASS PROPERTIESThe Spacecraft mass properties shall be fully compatible with the attitude control capabilityof the Spacecraft. The attitude control design shall account for mass property evolutions andmeasurement inaccuracies, physical build uncertainties, and propellant loading and locationuncertainties. Final mass properties verification shall be performed immediately prior toSpacecraft shipment to the Launch Site. The addition or removal of hardware, liquids, orgases, after the final mass properties measurements shall be logged and verified forcompatibility with the attitude control capability of the Spacecraft. A final mass properties /attitude control report shall be delivered to the Customer for approval prior to launchincorporating data obtained after propellant filling and final Spacecraft preparations to verifyand confirm that the mass properties fall within previously analyzed attitude controlcapabilities during the entire mission life.




6.4.9 AUTONOMYThe satellite shall be capable of maintaining continuous satisfactory operation without groundcontact for minimum 48 hours period throughout its operational lifetime, assuming no failurecondition.

The satellite shall be capable of surviving, without ground contact, any 48 hours periodthroughout its operational lifetime, including any single failure condition, without sufferingfrom irrevocable consequences.

6.4.10 CONTROL ELECTRONICS FAULT PROTECTIONADCS shall ensure that a single event upset in the digital circuitry, memory ormicroprocessors, does not affect the performance. Extensive, continuously operatingdiagnostic and correcting features shall be included in processor systems, shall have a watchdog timer and in the event of it being triggered, shall attempt automatic recovery by a softreset/switch over to the redundant system. The status of these faults shall be provided intelemetry. ADCS shall have the capability for a ground command initiated test mode toperform fault detection of hardware/firmware functions. The ADCS design shall ensure thesoftware shall not be in an infinite loop in the event of any interface failure.

ADCS software shall have provision to uplink a software patch and execute it as either onetime option or periodically as required.

ADCS software shall facilitate monitoring of internal parameters which are normally notavailable through telemetry for diagnosis.

6.4.11 PROPULSION INTERFACESThe ADCS shall generate and provide control and activation signals to the propulsionsubsystem in order to perform all attitude control and station-keeping maneuvers consistentwith the ADCS and propulsion subsystem requirements of this document. It shall not bepossible for the ADCS to operate thrusters in any automatic on board mode of operationoutside the allowable duty cycle domain defined by the manufacturer. ADCS shall haveprovision to terminate thrusters firing in the event of unintended continuous thruster firing orpropellant leakage.

6.4.12 TELEMETRY INTERFACESSufficient telemetry data shall be provided to verify the response from the ADCS elements todetermine the health of the subsystem. This shall include the following, as applicable

a. The state of ADCS, all sensors and actuators including their outputsb. The controller outputs and the related control parametersc. The faults occurred on board including the automatic reconfigurationd. The cumulative thruster firing duration for each thruster




6.4.13 SPECIFIC ADCS TEST:Each attitude and control subsystem shall have been subjected to tests which comprisesimulations of all operational modes of the subsystem and in addition for qualification allinter-mode transitions that are expected to occur during the life of the satellite. Thesesimulations shall preferably use a minimum three degree of freedom servo table supportingthe ADCS equipment.

6.4.14 SATELLITE ON BOARD PROCESSOR MEMORY DATAEach address of the on board processor memory shall have the following information: itsphysical meaning, its routing and use in the control logic, its default value at turn on and itsrecommended operating value. For avoidance of doubt, the on board processor means all theunits on board the Spacecraft that have software and/or firmware that is accessible bytelemetry.

6.4.15 ADCS BLOCK DIAGRAMSContractor shall deliver block diagrams of the subsystems, their interfaces, in sufficient levelof detail to allow autonomous creation and validation of satellite procedures by Customer, aswell as troubleshooting during orbit anomalies and understanding of the role of anyparameter that can be accessed in telemetry. Block diagrams shall provide in-depthdescription of full end-to-end processing for each on-board algorithm, including filters andtransfer functions definitions, gains default values, activation period and conditions,interaction with other algorithms, and telemetry and command interface. Contractor shallprovide the Hardware-Software interface and interaction analysis. Contractor shall deliver thecomplete memory map of all on-board processor software. Each entry of the memory mapshall contain the name, address, type, default value, unit and summary meaning of theassociated software parameter. Contractor shall deliver detailed information concerning anyother software tools necessary for satellite monitoring and control from launch vehicleseparation to the end of operational life in both normal operation and back-up modes

6.4.16 ATTITUDE CONTROL AND DETERMINATION DATAA system schematics drawing indicating sensors and actuators, locations, orientations andfield of view shall be provided. Calibration values, biases and trends for sensors andactuators, Nominal and extreme values of control subsystem parameters shall be provided intabular form. In addition, all parameters and coefficients loaded either on the on-boardcomputer or on the attitude determination sensors hardware shall be provided.

6.4.17 FUEL SLOSHING DATAContractor shall provide to the Launch Service Contractor all spacecraft sloshing datanecessary to perform the combined Launch Vehicle-Spacecraft Dynamic analysis.




7 PROPULSION SYSTEM7.1 SUBSYSTEM REQUIREMENTSThe propulsion subsystem, in conjunction with other spacecraft equipment, shall perform allrequired orbit rising during GTO, attitude control and orbit maneuvers including transitionfrom drift orbit to operational orbit within the limits.

The subsystem design shall be compatible with major commercially available launchvehicles. After satellite separation from the launch vehicle the three-axis stabilized modefunctions shall be performed until the end of spacecraft’s useful life.

The contractor can propose a chemical or electrical or a mix of propulsion techniques toachieve the overall mission requirements. Detailed analyses have to be provided onreliability, contamination on Spacecraft, Power margins, disturbance on control system,Operational constraints if any etc.

The contractor shall provide a detailed plan for orbit circularization, orbit inclinationcorrection, attitude control and spacecraft repositioning functions. The redundancyphilosophy for all the operations shall also be specified.

Restriction if any for transportation of spacecraft after fuel filling shall be clearly brought out.These restrictions shall be in line with the launcher requirements.

7.2 OPERATIONAL REQUIREMENTS7.2.1 RELIABILITYThe assessed probability of the propulsion system and equipment meeting all functionalrequirements in the specified environments shall meet the life of 15 years in orbit.

7.2.2 SAFETYAll inherently hazardous elements of the equipment shall be designed and manufactured tomeet the design safety requirements as specified in the Product Assurance Plan. Personneland equipment safety shall be a prime consideration in the testing, handling andtransportation of all subsystem equipment per Product Assurance Requirements forSubcontractors.

7.2.3 OPERATIONAL LIFETIMEThe total Subsystem shall be able to meet all the requirements of these Specifications for 15years of operation in orbit with a margin of minimum one year for performance dispersions,except for the LAM, pressurant tanks and associated lines, for which the operational life shallbe one month after final pressurization in orbit.




7.2.4 STORAGE LIFEThe Propulsion System shall have a minimum storage life on ground of 4 years withoutmaintenance or refurbishment under protected environment.

7.2.5 MAINTAINABILITYThe manufacturing and integration shall not restrict the modularity and the accessibility ofany assembly or equipment of the subsystem allowed by the design. Replacement of failedpart shall be possible until launch.

7.2.6 GROUND OPERATIONSEnvironments experienced during equipment fabrication, transportation and storage shall becontrolled so as to be significantly less severe than the environments specified herein.

7.2.7 LAUNCHDuring launch the propulsion components will be exposed to the following mechanicalenvironmental conditions:

a. Linear acceleration.b. Sine vibration.c. Random vibration.d. Acoustic noise.e. Shocks.

The Propulsion System shall withstand mechanical environment conditions generated by thesatellite structure induced by the launcher on the structure, as defined in the launchers vehicleuser's manual.

7.2.8 TRANSPORTABILITYThe Subsystem, when integrated, shall be transportable by air, ship and road when packagedin accordance with the Statement of Work.

7.2.9 FILLING AND DRAININGThe filling and draining operation shall be possible at spacecraft level with the satellite in thevertical position. The pressurization of the propulsion system shall be possible, especiallyafter filling with the satellite in the vertical position. Drying operation shall be possible atspacecraft level in any position with required Fill and Drain valves accessible.




7.3 DESIGN AND CONSTRUCTION7.3.1 PARTS, MATERIALS AND PROCESSESThe selection, documentation, qualification, approval and control of all parts, materials andprocesses shall be in accordance with the requirements specified in the Schedule-D. Apartfrom this, some of the specific propulsion system requirements are mentioned below.

7.3.1.1 LUBRICANTSNo lubricant shall be used except where use of KRYTOX 240 AC (Devolatilized) isspecifically required.

7.3.1.2 DISSIMILAR METALSWhere it is impossible to avoid dissimilar metals in direct contact with each other or theirexposure to the same electrolyte, coating, plating or otherwise adequately protecting one orboth surfaces shall provide suitable protection.

7.3.1.3 FUNGUS RESISTANCEMaterials, which are nutrients for fungi, shall not be used.

7.3.1.4 COMPATIBILITY WITH CORROSIVE MATERIALSAll materials, subject to exposure to the propellants, ground test liquids or gas, cleaningmaterials or their fumes, or combustion products shall be compatible. The type of leadwires/coils to be specified in confirmation with MIL Standard.

7.3.2 MECHANICAL DESIGNUnpressurized propulsion equipment shall be vented to accommodate the specifiedbarometric pressure rates of change for both decreasing and increasing pressure.

7.3.3 THERMAL DESIGN DESCRIPTIONThe heat capacity, heat dissipation for operating and non-operating conditions and heat fluxfrom the equipment to the structure surface shall be listed in the Interface Data Sheets of thepropulsion equipment.

8 MECHANISMS AND PYROTECHNICS8.1 DESIGN REQUIREMENTSAll mechanisms shall employ designs qualified for the envelope of ground handling, testing,storage and flight environments. All mechanisms shall be capable of being actuated in groundtests at Spacecraft level, in either flight-like conditions or with solar wings and reflectorsremoved where necessary.




The Torque/Force Ratio is defined as the ratio between the Available Torque/Force of themotorizing element (i.e. motor, spring, etc…) and the Required Torque/force to move theappendages. The available torque consists of the worst-case torque available from themotorizing element. The Required Torque/Force consists of the worst-case loads, internal andexternal, to the mechanism (including frictional losses, external harness, appendages inertia,Spacecraft motion induced torques, etc) that are necessary to be overcome, in order to drivethe appendages. Numerically, the Torque Margin is the Torque Ratio minus one. Allmechanisms shall provide torque/force margin of at least 200 percent (3:1 torque ratio)relative to the worst case in-orbit requirements, including thermal effects, aging, mechanicalmisalignments, manufacturing tolerances and wear throughout the Design Life. The aboverequirement shall be satisfied from any position in the range of motion, under the assumptionof zero initial kinetic energy. When the motorizing element is a stepper motor, theassessment of the motorization margin shall use the same step frequency and steppingmethod as the Spacecraft motor driving electronics.

Accurate verification of deployment performance on ground shall be possible, including forthose devices (solar wings, reflectors), where 2- or 3-dimensional deployment is required.There shall be no thermal constraints to the operation of any mechanism at any time duringthe Operational Life.

All mechanisms shall be either fully redundant, or electrically redundant. Spring redundancyis considered compliant with this requirement. All mechanisms shall be designed to operateover at least 1.5 times the number of cycles enveloped by the total of ground tests and DesignLife, or 50 cycles, whichever is greater. No mechanism shall exhibit an increase of theresistive torque of 50% or more from the initial value over the Design Life.

a) Rotating devices requirementsRotating lubricated devices shall be capable of supplying lubricant for at least 1.5 times theamount required through Design Life. Their design shall provide for management of debrisgenerated within the unit.

For equipment utilizing slip ring and brush devices for power transfer, the Contractor shalleliminate mechanisms by which shorting may be encountered within the device.The Contractor shall, in addition, provide for management of debris generated within suchdevices.

The mechanical design shall incorporate provisions to minimize the loads to which ballbearing suspension systems are subjected during launch. No events of bearing assemblyinstability are acceptable. The bearing design shall be shown analytically and by test to havebearing assembly stability margin for dimensional variations of manufactured parts and forlubricant conditions expected throughout all mission phases.




With the exception of bearings and final output gears, all mechanisms elements includingtheir control electronics shall be fully redundant.

b) Deployment and Pointing Mechanism RequirementsDeployment mechanisms shall provide positive latching features, except where themechanism holding torque is designed to be at least three times the overturning torque.

Reflector pointing mechanisms shall be provided for all deployable reflectors. Reflectorpointing mechanisms shall have a minimum step size of 0.01 degrees in the RF bore sight.The pointing capability of the RF boresight shall be ± 0.2 degrees in both pitch and roll fromthe nominal position, with exception of the steerable beam, that shall be able to coveranywhere over the visible earth. The Spacecraft shall be designed to with no RF self-blockages inside the above-specified pointing zones.

The design of mechanism shall take into account the effects of all ground handling andtesting environments. Stowing pins or deployment devices shall be accessible withoutdisturbing the relative position of the mechanisms.

The status and/or the position of mechanisms (e.g. running, deployed position) should betelemetered without reliance on the observation of secondary effects.

8.2 PYROTECHNIC DEVICESAll pyrotechnic devices shall be fully redundant, including fire command paths, electricalcircuit, heating element, power supply, initiators and cutters/blades/releasers. The designshall ensure actuation by either initiator. Squibs shall be capable of being firedsimultaneously.

All pyrotechnic electrical harness connections shall be designed to preclude the connection ofany other pyrotechnic device other than the one intended. All actuators and pyrotechnicdevices shall be flight-proven designs. All elements and interfaces in the pyrotechnicsubsystem (comprising the pyrotechnic chains, on-board computers, launch operationequipment, ground support and test equipment and all software associated with pyrotechnicfunctions) shall be demonstrated to meet the Spacecraft mission requirements. Preferenceshall be given to designs which minimize shock loads.

All Spacecraft ordnance shall comply with the safety requirements imposed by theCompatible Launch Vehicle agencies.

All separation or release devices shall be designed such that adequate margin exists for therelease, without fragmentation of any part of the release system, and shall have appropriatecatchers to constrain debris. The design shall include seals to preclude Spacecraft




contamination. Any sealing system used to prevent contamination shall be demonstrated to beeffective for the mission requirements.If the propulsion subsystem incorporates pyrotechnically actuated valves, full redundancy isrequired for each valve (series in case of normally open and parallel in case of normallyclosed valves).

In case that the design of pyrotechnically actuated valves uses seal rings for sealing of thepyrocharge compartment against propellant, pressurant or outside, long-term integrity of theassembly in presence of seal degradation shall be demonstrated.

The pyrotechnic actuator circuit shall be protected and remain fully operational with anypost-firing squib resistance ranging from open to short circuit. In addition, the design shallensure that the power subsystem is not overloaded before, during or after the actuation of anypyrotechnic device operation.

All pyrotechnic devices shall be defined as critical items and the design, manufacturing andtest requirements shall be treated according to Schedule D.

9 ORBITAL OPERATIONS9.1 TRANSFER ORBIT

The orbit raising strategy has to be clearly given along with the required justificationcovering adequate visibility, performance deviations if any, including propellant budget andthe expected life projection;

TT&C, Sun aspect angle, power, thermal and attitude control constraints shall be compatiblewith transfer orbit mission duration up to 15 days and a launch window in line with therequirements of the selected launchers

The circularization strategy shall maintain a physical separation of at least 70 km with anygeostationary satellite;

The mission planning shall tolerate the 3-sigma dispersions of the Spacecraft injection and ofthe LAM burns sequence (if any).

9.1.1 ATTITUDE ERROR BUDGETThe attitude error is defined as the angle between the target and assessed thrust vector of amaneuver. The total attitude error for Apogee/Perigee boost maneuvers shall not exceed 3deg (3-sigma).

The attitude error budget shall include, as a minimum:




1. The attitude determination error;

2. The attitude re-pointing error in the case the Spacecraft design requires an attitude re-pointing before the maneuver;

3. The random variation of the inertial direction of the actual thrust vector during themaneuver;

4. The LAM misalignment

9.1.2 TRACKING NETWORKThe tracking network shall provide geometric visibility of at least 90 % of the geostationaryring.

9.2 STATION-KEEPING9.2.1 TERMINOLOGY1. Timing error: defines error in timing of total impulsive Delta-V achieved compared to

the timing of the equivalent impulsive Delta-V targeted.

2. Station-keeping cycle: defines a repetitive, predefined cycle of station-keepingmaneuvers starting with a drift/eccentricity control maneuver and followed by one or aseries of inclination or inclination/eccentricity control maneuvers and with separateinterleaved or combined momentum control maneuvers.

3. Co-location, if applicable with existing customer satellites, maintenance strategy has tobe defined and to be considered for planning SK.

4. Maneuver predictability: is defined as the accuracy with which the actual delta-Vgenerated by a maneuver can be predicted a priori, using available telemetry data and thePropulsion System model software.

5. Delta-V error: is defined as the discrepancy between the realized and predicted delta-V.

6. High thrust propulsion: defines a propulsion system that allows the generation of aminimum delta-V of 1 m/sec within 2 hours.

7. Low thrust propulsion: defines a propulsion system which does not fulfill the high thrustpropulsion criteria

8. Hybrid Inclination/eccentricity control: refers to a propulsion system using:a. low thrust thrusters generating a radial coupling larger than 3% of the normal




component during an inclination maneuver for inclination control and partialeccentricity control in combination with

b. high thrust thrusters allowing to control longitude and eccentricity

9. All low thrust propulsion: refers to a propulsion system using exclusively low thrustpropulsion for station-keeping

9.2.2 REQUIREMENTSThe Spacecraft design shall allow the Spacecraft location to be maintained within 0.05 degreeNorth/South and 0.05 degree East/West, with the Spacecraft operated at an orbital positionand for a station keeping cycle of 14 days. The pointing error budget shall however supposethat the Spacecraft will be controlled within 0.10 degree Pitch, 0.10 degree Roll and 0.2degree Yaw. The Spacecraft will be co-located with several other satellites which will movearound the Spacecraft. A minimum separation of 10KM between the Spacecraft and theother satellites (Emissions and Collocation) will be maintained.

The Spacecraft design shall allow the execution of all the maneuvers required for control ofdrift, eccentricity and inclination over a specified station keeping cycle, according to thespecifications of the following paragraphs. These specifications must hold for the followingcases:

a. for any day of the year (including the eclipse season) without reducing the PropellantLife;

b. for the primary maneuver modes;c. for the alternate primary or backup maneuver modes recommended to minimize the

propellant consumption or to cope with Spacecraft design or maneuver date/timeconstraints;

d. Considering angular momentum de-saturation requirements corresponding to the worstcase external torques during the entire mission.

The duration of the specified station keeping cycle is fourteen (14) days.

It is desirable to have a Spacecraft design that offers the capability to receive an upload of allmaneuver parameters required for one 14 day station keeping cycle, it shall be possible to setthe parameters for each maneuver individually. After a maneuver, or series of maneuvers, isuploaded, it shall be possible to modify any parameter of an individual maneuver up to sixty(60) minutes prior to start of thruster operation.

In case the Spacecraft uses low thrust propulsion, the Spacecraft design shall allow theSpacecraft location to be maintained within ± 0.05° North/South and ± 0.02° East/West for astation keeping cycle of 7 days. It shall be possible to set the parameters for each maneuverindividually. After a maneuver, or series of maneuvers, is uploaded, it shall be possible to




modify any parameter of an individual maneuver up to sixty (60) minutes prior to start ofthruster operation.

9.2.3 DRIFT AND ECCENTRICITY CONTROL MANEUVERSThe propulsion subsystem shall be able to support all drift and eccentricity controlmaneuvers. The contractor shall propose a detailed plan depending upon the selectedpropulsion system.

9.2.4 INCLINATION CONTROL MANEUVERSThe propulsion subsystem shall be able to support all inclination control maneuvers. Thecontractor shall propose a detailed plan depending upon the selected propulsion system.

9.2.5 HYBRID INCLINATION/ECCENTRICITY CONTROL MANEUVERSThe propulsion subsystem shall be able to support hybrid inclination and eccentricity controlmaneuvers (if required). The contractor shall propose a detailed plan depending upon theselected propulsion system.

9.3 ADDITIONAL REQUIREMENTS9.3.1 SPACECRAFT LONGITUDE RELOCATIONThe station keeping specifications in section 9.2 shall continue to be satisfied in case theSpacecraft is re-located to any longitude inside the interval specified in section, withoutmodification of the antenna boresight.

The Spacecraft design shall support a minimum of two (2) relocations with a minimum driftvelocity of 2.0deg/day during Design Life. It shall be possible to achieve this relocation ratewithin 2 days.

9.3.2 RANGING SYSTEM REQUIREMENTSDuring the transfer orbit, in-orbit tests and in case of Spacecraft control via the omni antenna,the net (total error budget) slant range accuracy shall be within 15 meters (3 sigma).

9.3.3 MAXIMUM DELTA-V FOR ATTITUDE RECOVERY MANEUVERSThe maximum Delta-V components required for attitude recovery maneuvers (Sun or Earthacquisition or a momentum wheel switch) for a Spacecraft with no more than one failure,shall be as listed below:

a. Delta-V radial < 0.25 m/sb. Delta-V tangential < 0.10 m/sc. Delta-V normal < 0.50 m/s




9.3.4 DISPOSAL ORBIT AND PASSIVATIONThe Spacecraft shall be capable of being put in a disposal orbit and being passivated in acontrolled way at end of Propellant Life. The disposal orbit shall have an apogee and aperigee of at least 350 km above geo-stationary altitude.

9.4 PROPELLANT BUDGETThe Contractor shall clearly derive and substantiate all the numbers that build up thepropellant budget.

In particular, the propellant budget shall include sufficient amount of propellant andpressurant to make allowance for all causes of inefficiency of the operation, including but notlimited to the following:

a. thruster cant and misalignment;b. plume impingement/ drag;c. orbit and attitude determination uncertainties;d. constraints imposed by the other subsystems of the Spacecraft during transfer and

geostationary orbits;e. mission profile where firings occur at non-optimum times or locations;f. cross-coupling effects associated with orbit and attitude control maneuvers;g. leakage and solubility;h. thruster selection and maneuver strategies, including backup modesi. delta velocity uncertainties due to thruster performance dispersion;j. thruster firing modes (duty cycles) and blow-down effects;k. mixture ratio effects;l. 3 sigma low performance of the launch vehicle and the Spacecraft on-board thrusters and

LAM.

The propellant budget shall account for the following:

a. Orbital position requiring the larger Delta-V requirements for longitude control in therange specified;

b. An additional Delta V requirement for In-Orbit-testing of 0.20 m/s for East/West controlmaneuvers and 6.0 m/sec for North/South control maneuvers;

c. A re-location of 2.0 deg/day (total of 11.4 m/sec) after the completion of the In-Orbittesting in order to re-locate the Spacecraft to the final position;

d. One re-location maneuver of 1.0 deg/day (total of 5.7 m/sec) at mid-Propellant Life;e. A propellant reserve shall be included at end of Propellant Life for Spacecraft retirement

to enable an increase of the semi-major axis by 350 km with circularization of the orbit(total of 10.9 m/sec).




10 DESIGN AND ANALYSESThe Contractor is responsible for performing all design and analyses activities in theprogramme including preparation of design analyses, reports as well as planning andorganizing of a series of design reviews.

10.1 DESIGN ANALYSES AND STUDY REPORTSThe Contractor shall perform comprehensive design analyses and studies whichdemonstrate the integrity of the satellite and compliance with its specifications, includinginterface specifications. Margins shall be demonstrated and quantified for all modes ofoperations and environmental exposures from manufacture and tests on the ground throughto the end of the operational lifetime.

10.1.1 REQUIREMENTS:The design analyses and studies shall address the following topics:

a) definition of the functional, performance, operational, design and interface andenvironmental specifications and as well the interactions between the variousequipment and subsystems or interfaces;

b) Design description and justification. The justification of the proposed design shall bebased on analyses and studies in all appropriate disciplines. The analyses shall besupported by the results of appropriate development model tests. A design report shallbe produced per equipment for each equipment and subsystem.

Sufficient details shall be provided by the Contractor to allow complete verification of allanalyses and studies by ISRO. In addition, the Contractor shall provide in thedocumentation any explanatory test data, drawings, circuit diagrams, references or anyother material that will be beneficial to the review of the information provided.

Analyses shall be updated by the Contractor as new information becomes available during theContract period. The report on the updated analysis shall describe any differences from theoriginal report. Background information contained in the original report need not be repeatedunless sufficient change has occurred such that ISRO requires the new report to be self-contained.

The general design analyses and study report requirements set forth in this section shall haveprecedence over the detailed description of design analysis requirements presented insubsequent sections. The detailed design analyses requirements illustrate the typical scope ofwhat is anticipated to be included.

Nevertheless, any additional analysis determined to be necessary to satisfy these generalrequirements shall be considered to be within scope.




10.1.2 ANALYSES DATAThe Contractor shall provide at ISRO premises all the data which is necessary to repeat atISRO premises the analyses performed to validate the design.

10.2 SYSTEM LEVEL ANALYSESThe following defines the minimum requirements for satellite system level analyses to beperformed by the Contractor. These analyses shall be documented in the satellite systemdesign report.

At satellite level, the design description shall address the satellite design from an overallsystem level. This shall include the major satellite/subsystem trade-offs (if any), the overallsatellite configurations (mechanical and electrical), and the satellite interfaces. Theseinterfaces shall include those of the applicable launch vehicle(s), telemetry and commandlists, mission analyses with budgets, and the satellite environments (vibration and shock,thermal constraints, radiation survivability, etc.). The inter-relationships of the varioussubsystems are also to be addressed in the form of power summaries/profiles, massproperties, equipment layouts, integration sequence, system block diagrams, electromagneticcompatibility (EMC) plans and electrical interface control requirements. The equipmentlayout optimisation shall be performed with due consideration of the maintainability andaccessibility requirements.

10.2.1 DYNAMIC ANALYSESA dynamic analysis of the satellite shall be performed by the Contractor for the variousdynamic conditions during the different mission phases. This analysis shall include as aminimum the effects of propellant slosh, centre of gravity shifts and flexure of non-rigidstructures. The analysis shall address the full range of propellant loading and the associateddynamics.

Satellite separation dynamics analyses for the specified launch vehicle(s) shall beperformed.

10.2.2 PROPELLANT BUDGET ANALYSESThe Contractor shall analyse the propellant requirements of the satellite for all phases of themission and provide a propellant budget identifying the necessary propellant margins, basedon worse case or adverse 3 sigma performance.

The budget shall include quantities for contingency margin, thruster misalignment, thrusterefficiency losses, subsystem residuals, leakage, and dispersions from firings. It shall alsoinclude provision for one pattern mapping test campaign for each antenna at the beginning oflife. It shall address all worst case combination of thrusters for operation of the propulsionsubsystem.




In addition it shall also include a provision for re-orbiting at the end of life and longituderepositioning.

The Contractor shall provide details of the performance prediction techniques used to predictrestrictions of operational lifetime due to propellant consumption and assessment ofremaining propellant on board.

10.2.3 MISSION ANALYSESThe Contractor shall perform detailed mission analyses as defined in Schedule F, the resultsof which shall be used in the propellant budget analyses and also to define the FlightOperations Plan which provides the detailed sequence of events for all in-orbit operations.

10.2.4 MASS PROPERTIES ANALYSESThe Contractor shall perform mass properties analyses which shall include, but not be limitedto, a detailed listing of the mass of each satellite unit, their relative locations and theircontributions to the satellite centre of mass location, and moment of inertia matrix. Theanalyses shall be performed for all mission phases and configurations and for all applicablelaunch vehicles selected by ISRO.

10.2.5 ELECTROMAGNETIC COMPATIBILITY (EMC) ANALYSESThe Contractor shall define the EMC specification to which all the satellite equipment mustcomply and perform analyses and tests to demonstrate that the satellite design will adequatelycope with radiated and conducted emissions and susceptibility as well as the magnetostaticenvironment.

The Contractor shall perform analyses to demonstrate radio-frequency compatibility of aISRO Ka band HTS collocated with other satellites.

The Contractor shall perform analyses to demonstrate that the satellite design complies withthe constraints imposed by the launch authorities.

10.2.6 ENVIRONMENTAL EFFECTS ANALYSESThe Contractor shall analyse the effects of the environmental exposures that the satellite willexperience prior to launch, during launch, up to placement in geostationary orbit andthroughout its operational life in orbit. These analyses shall, at least, include:

10.2.6.1 RADIATION EFFECTS ANALYSESThe Contractor shall evaluate the radiation levels to which each satellite unit will be exposedand implement any protection accordingly. The radiation effect analysis shall be sufficientlydetailed to demonstrate the correct operation of all electronic components and equipmentincluding Single Event Upset (SEU) and Single Event Transient (SET) immunity until the




end of operational lifetime after exposure to the worst expected case of radiation environmentduring transfer orbit, drift orbit and on-station operation.

10.2.6.2 SATELLITE CHARGING ANALYSESThe Contractor shall perform system level analyses to demonstrate that the satellite will beimmune to worst case environments for Electrostatic Discharge (ESD). The analyses shallinclude a surface and bulk charging analysis and transient response predictions of the satelliteand critical subsystems. The analyses shall be supplemented with grounding schemes andrequirements for exterior and interior satellite surfaces and shall describe measures whichwill be employed to avoid or reduce the effects of discharges between different parts of thesatellite in-orbit.

The analyses and the results of ESD testing shall demonstrate the adequacy of the ESDimmunity of the satellite design for the operational lifetime.

10.2.6.3 SATELLITE CONTAMINATION ANALYSESThe Contractor shall perform analyses and prepare plans to avoid or control any external andinternal contamination during all phases of manufacturing, integration, testing, storage,transportation, launch and operation of the satellite after separation from the launch vehicleand throughout the mission. Contamination sources shall include, but not be limited to,outgassing products from the satellite, thruster plume impingement and variation ofcleanliness induced by external sources.

10.2.6.4 SATELLITE DEPRESSURISATION ANALYSESThe Contractor shall perform analyses to demonstrate that the satellite will withstand theworst-case depressurisation environment as specified by the launch vehicle agencies.

10.2.7 PAYLOAD ANALYSESThe Contractor shall provide detailed analyses to demonstrate the adequacy of the design andconfiguration of the Payload and the conformance of its performance to the communicationrequirements in all modes of operation.

10.2.7.1 COMMUNICATION LINK ANALYSESDetailed information and link analyses shall be provided to allow ISRO to verify theperformance of the Payload for every requirement specified in Schedule B, TechnicalSpecifications Requirement.

10.2.7.2 ANTENNA POINTING ERROR ANALYSESThe Contractor shall perform Antenna Pointing Error Analyses, taking into account allpotential contributions of disturbances and errors, to define the beam pointing errors to beused in the coverage related performance analyses. This budget shall present the contribution




of all the various factors for each antenna. A separate antenna pointing error analysis andbudget shall be provided for the in-orbit antenna mapping test.

Based on the above analyses the Contractor shall determine the worst case, lifetime and dailypointing error ellipses for each antenna. The lifetime error ellipse shall be calculated usinglong term, diurnal and short term error contributions. The daily error ellipse shall becalculated using diurnal and short term error contributions.

10.2.8 FDIR RESPONSE AND OBSERVABILITYThe Contractor shall perform an analysis of the satellite FDIR response to each failure modeidentified in the FMECA. The analysis will demonstrate that the recovery actions areexecuted correctly, and identify how the failure and how the recovery will be observable intelemetry.

10.3 SUBSYSTEM LEVEL ANALYSES10.3.1 TELEMETRY, TELECOMMAND AND RANGING ANALYSESThe Contractor shall provide analyses including TCR antenna patterns and link budgetsaccording to the ground station characteristics, worst case margins and interference levelsfrom the communications signals. This link data will be provided as an annex to theSatellite/Ground Station Interface Requirements. For phase lock loop receivers the noisebandwidth and carrier acquisition requirements shall be indicated. Telemetry and Commandchannel budgets, showing spare capacity shall be provided.

10.3.2 THERMAL ANALYSESThe Contractor shall perform complete and comprehensive thermal analyses including allappendages (e.g. antenna reflectors, solar arrays).

All modes of operation of the ISRO Ka band HTS shall be analysed including as a minimumstorage, pre-launch (including encapsulated conditions and on launch pad), launch, parkingorbit (if applicable), transfer orbit, perigee (if applicable) and apogee motor firing transients,drift orbits and acquisition/deployment sequences, beginning and end-of-life solstices andequinoxes, including transient conditions, plume heating and any safe modes.

The thermal effects of individual unit (e.g. high power isolator) or transponder channelfailures shall also be considered. Special care must be taken with analyses of interfaces withcritical satellite elements (e.g. antenna reflectors, solar arrays).

All design cases shall include thermal-optical performance degradation of thermal controlsurfaces over the satellite operational lifetime.




The thermal analyses shall be documented and submitted to ISRO as part of a design analysesreport. In addition the analyses shall describe the method defining the conduction andradiation exchange factors.

Moreover, in order to allow the launch vehicle authorities to carry out a coupled thermalanalysis, the Contractor shall provide a thermal model of the satellite, compliant with thespecification required by the launch vehicle authorities.

The Contractor shall provide temperature prediction for system level testing and flight modelsatellites which shall be based on the thermal balance test correlated model using worst casemodel parameters and environment for all mission phases.

10.3.3 STRUCTURAL ANALYSESThe Contractor shall perform analyses in which all major structural elements, includingantenna reflectors, antenna deployment mechanisms and solar arrays are investigated. Thedesign factors or margin of safety shall be defined and demonstrated by the Contractor,including design loads applied to the primary structure.

The satellite structure shall be analysed, both statistically and dynamically, to demonstratethat all structural elements perform adequately when subjected to anticipated load conditionsand the ability of critical structure elements to maintain required in-orbit dimensional stabilitywith the specified margins.

The compatibility of the critical structure elements with the attitude control subsystem shallbe demonstrated for transients created by thrusters or by deployment of the appendages or bythermal transients at the beginning and at the end of eclipse periods.

The mechanical stress analyses shall consider fatigue, creep, and stress relaxationcharacteristics. When critical, these characteristics shall be analysed in detail and comparedto applicable empirical data. If there is no empirical data, the Contractor shall conduct testsadequate to confirm the design.

The mathematical model shall be determined by a modal test and the model inaccuracy shallbe determined and applied on the results of the analyses.

The Contractor shall provide a reduced node structural model compliant with the formatrequired by the launch vehicle authorities for use in satellite/launcher coupled load analyses.The model will be delivered to ISRO prior to CDR.

A dynamic response analysis performed with the mathematical model of the satellite shalldemonstrate specified margins with respect to the launch environment and detection ofpotential problems during system level tests and coupled load analysis results.




10.3.4 ELECTRICAL POWER ANALYSESThe analyses to be provided by the satellite Contractor shall include, at least, the following:

A detailed description of the power subsystem with power required by each satellite unit, ineach of the possible modes of operation.

Assessments of solar array voltage-current characteristics under worst case conditions,including beginning and end of design life, equinox, solstice. The types of degradation effectsand the rates and method employed for calculating the power output shall be clearly definedand reviewed;

Power profile showing worst case daily, seasonal and lifetime variations in the current andvoltage requirements and the availability of power from the subsystem during all phases ofthe satellite mission;

Battery orbital operating conditions, battery management plans including charge control,charge period, times of reconditioning (if necessary), temperature effects and redundancyanalysis, and other data required to adequately support the battery design selected;

Electric power conditioning system including normal operation, redundancy design approach,transients, regulation, ripple, grounding philosophy, and other data needed to fully define alloperating modes of the system;

Analyses and test data shall be provided as appropriate to demonstrate that shadowing andhot spot failure modes and possible breakdown of insulation between solar cells and panelstructure or substrate have been accounted for in the design and that satellite propulsionsubsystem exhaust plumes or any launch vehicle related contamination shall not producedetrimental effects upon the array performance;

All circuits in the power subsystem shall have worst case stress and performance analysesperformed to adequately demonstrate performance and stability margins and shall be deratedon the basis of maximum voltage, current, temperature and power stress under all modes ofoperation experienced in orbit and during ground testing and at both extremes of operatingtemperature;

DC power budget analysis for all critical phases of satellite life including,at least, the initial and end of design life, equinox, solstice, parking andtransfer orbit condition.

Provision of all necessary analyses and operations procedures to predict the maximum safeuse of communications channels in excess of the specified EOL number of channels as a




function of satellite lifetime to assess solar array and battery power margins together with anyother satellite constraints.

10.3.5 ATTITUDE DETERMINATION AND CONTROL ANALYSES

A stability analysis for each control mode showing stability margins under nominal and worstcase system level inputs.

A performance analysis/simulation for each control mode with nominal and worst casesystem level inputs.

An analysis of the dynamic interaction between structural elements and the attitude controlsubsystems;

An analysis of dynamic behaviour in case of failures and anomaly conditions, identifyingrisks of entry into abnormally large attitude dynamics/flat spin condition and scoping therequirements for recovery.

An outage analysis showing safe modes with estimates of nominal and worst case times torecover normal pointing in accordance with the safe mode design;

An analysis of the impact of single failures on the operation of the safe modes;

An analysis of nutation control during all applicable mission phases;

Verification by simulation and analysis that all foreseen mode to mode transitions arepossible and dynamically smooth;

An analysis showing that all failures identified in the FMECA are covered by the FDIR shallbe provided.

10.3.6 PROPULSION ANALYSESThe Contractor shall perform a comprehensive set of functional, performance, design andinterface analyses and provide analytical support documentation.

a) Cleanliness requirements and control;b) Method and accuracy of subsystem and equipment leak check;c) Subsystem thermal analysis for all mission phases;d) Pressure budget;e) Feed system pressure drops acceptability;f) Thermodynamic description of pressurisation system behaviour during apogee

manoeuvres;




g) Structural/strength and facture mechanics analyses (where appropriate) for criticalequipment bearing pressure loads (e.g. Propellant and pressurant Tanks);

h) Thruster performance analysis for all mission phases (equipment andsystem level);

i) Detailed thruster performance data;j) Thruster integrity and performance assessment in case of gas ingestion;k) Thruster plume model and impingement effect analysis;l) Propellant sloshing analysis for all mission phases;m) Propellant tanks operational design constraints;n) Propellant tank expulsion efficiency and residual analysis;o) Justification of the maximum tank fill ratio;p) Analysis of subsystem component performance uncertainties on propellant

budget;q) Analysis of propellant mass measurement method accuracy;r) End of life propellant management analysis and strategy;s) Liquid and gas pressure data during subsystem activation;t) Attitude control subsystem requirements for the propulsion subsystem;u) Liquid and gas pressure transient analysis during subsystem activation and component

operation interaction analysis;v) Analysis of material compatibility with propellant simulants and pressurants.

ISRO


ISRO Ka BAND HTS

SCHEDULE C

ASSEMBLY, INTEGRATION, TESTINGAND SATELLITE STORAGE PLAN

ISRO/ISAC/HTS/RFP/001/06VER-01 SCHEDULE - C

ASSEMBLY, INTEGRATION, TESTING & SATELLITE STORAGE PLAN

December 2015 Page C-1ISRO CONFIDENTIAL PROPRIETARY

TABLE OF CONTENTS

SCOPE ........................................................................................................................................31 SUBSYSTEM LEVEL TESTING........................................................................................ 32 SYSTEM LEVEL INTEGRATION TESTS.......................................................................... 32.1 AIT SEQUENCE ................................................................................................................3

2.2 GROUND SUPPORT EQUIPMENTS ....................................................................................4

2.3 ADDITIONAL TEST REQUIREMENTS................................................................................4

2.3.1 ELECTROMAGNETIC COMPATIBILITY TEST ...................................................................4

2.3.2 ESD TESTS.......................................................................................................................4

2.3.3 ALIGNMENTS ..................................................................................................................5

2.3.4 INTERFACE VERIFICATION TESTS ..................................................................................5

2.3.5 SPACECRAFT RF, BASEBAND AND SOFTWARE COMPATIBILITY TESTING ....................5

3 SPACECRAFT HARDWARE TRANSPORTATION AND HANDLING................................. 63.1 SAFETY OF ALL OPERATIONS .........................................................................................6

3.2 SAFETY OF MECHANICAL OPERATIONS..........................................................................6

3.3 SAFETY OF ELECTRICAL OPERATIONS............................................................................ 7

3.4 SAFETY OF THERMAL VACUUM OPERATIONS ................................................................7

4 SPACECRAFT STORAGE PLAN ......................................................................................7




SCOPEThis Schedule deals with the Subsystem level and System level tests on spacecraft. Thesection also details information about safe handling of spacecraft during AIT build and alsoactivities pertaining to transportation of spacecraft to the Launch Site, post launch operationsincluding the storage plans.

1 SUBSYSTEM LEVEL TESTINGSubsystem development tests shall be conducted to fully justify, as a complement to designanalysis, the adequacy of the subsystem design. The qualification tests shall demonstratedesign margins built into the subsystem.

A subsystem integrated at satellite level should have successfully passed its integration testsprior to Satellite performance test. The acceptance tests shall demonstrate that the subsystemhas the specified performance in all modes of operation and can survive environmentalexposure. All subsystem level performance specifications shall be met before, during andafter the acceptance test. Detailed subsystem level test plans, test procedures, along withacceptance plans shall be provided by contractor and will be applicable subject to review andacceptance by ISRO.

2 SYSTEM LEVEL INTEGRATION TESTS2.1 AIT SEQUENCEAfter successful completion, evaluation and acceptance of the subsystems, the same shall beoffered to AIT for the next set of operations on the spacecraft. The Spacecraft shall besubjected to a system Level build and test program that includes all the tests required to assessthe performance of the Spacecraft and its subsystems, in the various test phases as part ofAIT sequence. A typical sequence can involve the following major phases of AIT on thespacecraft. This is only an indicative sequence; however the contractor shall propose theirAIT sequence, detailed AIT plan and procedures. The various test parameters, environmentalconditions, spacecraft configuration, test setup, detailed test matrices for each test phase withacceptance/ rejection criterions shall be submitted. There shall be a test readiness review,before start of each major phase and test results review after end of each major test phasebefore proceeding to the next phase. Applicable documents, safety certificates and equipmentcalibration certificates shall also be made available before start of each test phase.

The typical major AIT phases are as follows:

a. Incoming inspection and acceptance of each subsystemb. Disassembled mode Test Phasec. Assembled Mode test Phased. Thermo Vacuum (TVAC) Test Phasee. Thermal Balance Test Phase




f. Deployment test phaseg. Dynamic Test Phase (Vibration and Acoustic test phase)h. Post Dynamic Test Phasei. Compact Range Test Phasej. Pre Shipment Test Phasek. Launch Base Test Phase

2.2 GROUND SUPPORT EQUIPMENTSContractor shall provide all the Ground Support Equipments (GSE) including MechanicalGround Support equipments (MGSE), Electrical Ground Support equipments (EGSE) alongwith test equipments and test facilities needed for carrying out IST operations on thespacecraft. The contractor shall implement operating procedures to ensure that all Spacecrafthardware is not overstressed during ground handling. Contractor shall involve ISRO teamduring evolution of architecture, realization and subsequent Test & Evaluation of the GroundSupport Equipment (GSE) required for Spacecraft test and build activities. Designdocuments, User Manual and Test & Evaluation documents shall be made available tocustomer for review & approval of GSE. Both hard and soft copies shall be provided tocustomer. Customer shall have technical visibility into the GSE and operating procedures butshall not take delivery of the GSE.

Contractor shall also make available Spacecraft GSE to be used during the pre-launch phasesof the mission and ensure that a sufficient amount of spare GSE equipment is available at thelaunch site. In particular, spares need to be available for all critical countdown equipment.

2.3 ADDITIONAL TEST REQUIREMENTSThe Contractor shall also carry out the following additional activities as part of the ISTprogramme.

2.3.1 ELECTROMAGNETIC COMPATIBILITY TESTThe fully assembled Spacecraft shall be submitted to a series of electromagnetic compatibilitytests. These tests shall include auto compatibility (immunity to RF generated by the satellite),as well as susceptibility tests. As a part of this test phase, the electromagnetic fields near theADCS sensors shall be measured to demonstrate they are within their predicted maximumvalue. All ADCS sensors shall be tested for radiated susceptibility as part of such test byillumination with test horns over the spectrum of Customer’s fleet payload frequencies.

2.3.2 ESD TESTSThe adequacy of the Spacecraft design to survive electrostatic discharge threats shall bedemonstrated by a series of tests and analysis to be performed in the qualification model. It isexpected that such demonstration have been already performed on the proposed design inprevious qualification programs.




In such case, Contractor shall demonstrate the deviations and departures from the original testarticle to the Spacecraft design are within the envelope of the qualified design. The ESDanalysis carried out for the spacecraft configuration, materials used for spacecraft realization,ESD mitigation techniques followed and analysis results are to be given by the contractor forevaluation.

The demonstration of ESD resistance by qualification tests shall also apply to any unit that isdeemed as particularly susceptible to ESD threat due to its location in the Spacecraft orparticularities in its design. The contractor shall describe the units that fit into this category,and describe the tests performed to demonstrate ESD resistance of such units.

The Spacecraft flight hardware shall not be submitted to ESD testing.

2.3.3 ALIGNMENTSContractor shall measure and align all mechanisms, sensors, antennas, and thrusters on theSpacecraft in (Assembled Mode) Initial IST and FIST. Data derived from these measurementsshall be compared and the differences between both sets shall be compared with therepeatability limits defined in the alignment budget. If any mechanism, sensor, antenna orthruster is removed or has its configuration altered after the post-environmental alignmentmeasurement, then the affected hardware shall have its alignment re-measured and re-adjusted. Contractor shall measure the alignment of all mechanisms, sensors, antennas, andthrusters on the Spacecraft immediately after the vibration tests, to ascertain any anomaly ordamage caused by the exposure to the mechanical environment.

2.3.4 INTERFACE VERIFICATION TESTSInterface verification tests shall be performed to verify RF compatibility of the GroundControl System (GCS) with the Spacecraft and to verify the electrical and mechanicalinterfaces between the Spacecraft, the launch vehicle adapter and separation system, and thelaunch vehicle.

2.3.5 SPACECRAFT RF, BASEBAND AND SOFTWARE COMPATIBILITYTESTING

Contractor shall perform testing of appropriate GCS and ground segment equipment with theSpacecraft as part of the Integrated System Test Phase (either at the Initial or at the FinalIntegrated Test Phase) to verify ground segment to spacecraft RF, baseband and softwarecompatibility. Customer shall support spacecraft-to-ground RF compatibility testing byloaning applicable ground control equipment and personnel to set up the equipment.




3 SPACECRAFT HARDWARE TRANSPORTATION ANDHANDLING

Contractor shall be responsible for the safe handling and transportation of all Spacecraft sub-assemblies and equipment. As a minimum, the Contractor shall implement the followingmeasures throughout the program:

3.1 SAFETY OF ALL OPERATIONSThe ground support equipment design shall preclude common control of different pieces ofground support hardware, i.e. ground support equipment shall have its unique control, toprevent unintended actuation of the wrong component of the support equipment.

3.2 SAFETY OF MECHANICAL OPERATIONSConcerning safety of mechanical operations performed on sub-assemblies and equipmentrequiring lift, transportation or handling for test and/or integration purposes:

1. Write individual procedures for every specific handling operation or transportationactivity of flight hardware.

2. Validate the procedures by dry runs with the individuals and the hardware assigned totheir execution under supervision of a responsible independent authority.

3. Whenever a move, transportation, tilt or lifting of the flight hardware is planned to beexecuted, Contractor shall conduct a meeting before its execution in the same shift,with the assigned personnel to go through the actions to be performed, undersupervision of the safe handling and transportation certified responsible in charge.

4. Contractor shall only utilize ground support equipment to transport, move, tilt or liftthe flight hardware that has security features to prevent its operation in an unintendedor unsafe mode and designed to alert operators through appropriate alarms.

5. Contractor shall only utilize ground support equipment to transport, move, tilt or liftthe flight hardware that permits the verification of the correct attachment of thehardware concerned by direct visual inspection, i.e., all bolts and attachment partsnecessary to secure the flight hardware shall be clearly visible.

6. Contractor shall include an accountability check for every ground support equipmentin all procedures.

7. Before attaching any device to the flight hardware, Contractor shall verify that ageingof the device has not degraded its properties, resulting in a hazardous situation.




8. Contractor shall post warnings in visible location in the area of activities, to make thepersonnel aware of the possible significant mishaps.

9. Contractor shall design the operations in such a way to allow the operator to be able todirectly watch the effects of the actions of such operator in complete safety.

3.3 SAFETY OF ELECTRICAL OPERATIONSConcerning safety of electrical operations, the following requirements apply to all flighthardware from unit to Spacecraft level:

1. Contractor shall only utilize to connect to flight hardware power supplies that haveadjustable voltages and current limits. Contractor shall adjust such limits to ensuremaximum protection in case of shorts or malfunction. Contractor shall ensure UPSpower from 2 sources to ensure that failure of one system will not cause unintendedswitch-OFF of the space craft and related GSE.

2. Contractor shall write procedures specific to every power supply to be connected tothe Spacecraft or individual units and subsystems, containing instructions to correctlyset voltages and safety trip-off devices, and execute such procedure before connectingthe supply to the equipment.

3.4 SAFETY OF THERMAL VACUUM OPERATIONSConcerning safety in thermal vacuum operations, the following requirements apply to allflight hardware from unit to Spacecraft level:

1. Contractor shall only utilize uninterrupted power supplies to power the flighthardware and the thermal vacuum chamber and thermal regulation systems.

2. Whenever testing with high power RF or high voltage conditions, Contractor shall putin place sensing and power cut off systems to automatically protect the flighthardware in case of anomalies that result in damaging reflected RF power levels orcorona arcing conditions.

a. Thermal vacuum facilities for high power RF equipment shall be equippedwith alarms and safeguards to detect and shutdown the Spacecraft hardware inthe event of vacuum leaks

4 SPACECRAFT STORAGE PLANIn case it is decided to keep the spacecraft in storage due to the conditions arising at any pointof time either at the place of development or at the place of launch of spacecraft, theresponsibility of keeping the same in storage shall lie with the contractor. The contractorshall submit a detailed proposal for the storage, management of spacecraft during storage,removal of spacecraft from storage, the tests and preparation to be carried out on spacecraft tomake it ready for next set of activities. The contractor may propose the cost implication forthe same on basis of monthly storage plan.

ISRO


ISRO Ka BAND HTS

SCHEDULE D

PRODUCT ASSURANCEREQUIREMENTS

ISRO/ISAC/HTS/RFP/001/06VER-01 SCHEDULE - D

PRODUCT ASSURANCE REQUIREMENTS

December 2015 Page D-1ISRO CONFIDENTIAL PROPRIETARY

TABLE OF CONTENTS

1 PROGRAM CONTENT ..................................................................................................... 32 PRODUCT ASSURANCE REQUIREMENTS ...................................................................... 33 SPACECRAFT PRODUCT ASSURANCE MANAGEMENT .................................................44 REVIEWS AND AUDIT ..................................................................................................... 45 EEE PARTS SELECTION & CONTROL ............................................................................ 56 MATERIALS & PROCESS CONTROL .............................................................................. 57 RELIABILITY ASSURANCE............................................................................................. 58 SOFTWARE QUALITY ASSURANCE ...............................................................................69 MODEL PHILOSOPHY AND TEST ................................................................................... 6




1 PROGRAM CONTENTContractor shall conduct, from program inception to delivery and in-orbit hand over of theSpacecraft, a Product Assurance Program (PAP) compliant with the provisions of this plan.The purpose of the PA program is to actively assure and verify that the design, constructionand test of the Spacecraft and all of its Subsystems, Units and Components fully meet allthe specifications throughout its Design Life.

This plan shall constitute the master planning and requirements document for the SpacecraftPAP, describing the methods of implementation of R&QA requirements and the qualitystandards to be applied. The PAP shall comprise:

a. A quality assurance program including software control activities;b. A reliability assurance program;c. A parts, materials, and processes control program.d. A configuration control program

Responsibility within the Contractor's organization for ensuring the implementation of thesedisciplines shall be assigned as prescribed by the Contractor's Program Management Plan.Contractor is also responsible for ensuring & passing on the PAP requirements to all levelsof subcontractors.

Detailed ISRO requirements shall be made available on ISRO web site. The online links tothis shall be provided separately. However major highlights are given in this schedule.

• The contractor shall ensure & demonstrate to customer that the spacecraft & itssubsystems are designed, procured, manufactured, tested and delivered inaccordance with the PA requirements/guidelines agreed upon.

• During the entire project realization period, from placement of order to delivery ofspacecraft, contractor shall provide unrestricted access to-design documents,fabrication labs/records, test results, NCs with associated deliberations etc. toresident/itinerant representative of ISRO

• Customer shall have the right to perform independent audits of the effectiveness andimplementation of the Contractor's product assurance program, at Contractor's,Subcontractors' and contractors' facilities at any time during the Spacecraft program

• Customer reserves the right to participate in reviews/take decisions on critical issues

2 PRODUCT ASSURANCE REQUIREMENTSThe contractor shall establish & verify/demonstrate that all subsystems/systems aredesigned, manufactured, tested, stored and delivered to meet the specified performancerequirements throughout their design life.The contractor shall generate a PA plan which will constitute a master plan & requirement




document.

The product assurance plan shall include• Spacecraft PA management• Configuration and data management• Subcontractor/contractor management/control• EEE Parts selection & control• Critical item control• Materials & Process/manufacturing control• Cleanliness and contamination control• Design reviews/audits• Reliability analyses• Software quality assurance• Model philosophy & test plan• Non-conformance control & management• Storage, maintenance & recovery plan

3 SPACECRAFT PRODUCT ASSURANCE MANAGEMENTIn order to establish that all the subsystems/systems are designed, manufactured, tested,stored and delivered to meet the specified performance requirements throughout theirdesign life,

• Contractor shall identify a full time PA manager for the project• Contractor shall create a Product Assurance setup/team with experts under overall

responsibility of PA manager• PA manager along with his team shall have the authority to ensure the

implementation of a reliable design based upon proven design practices• Ensure early & prompt detection and reporting of actual or potential deficiencies,

marginal quality/trends/conditions that could result in unsatisfactory performance.• He shall also be responsible to monitor & ensure that all subcontractors comply with

the PA requirements/specifications• He shall have the authority to halt the activities, any time & anywhere, that may

violate the PA requirements/jeopardize the reliability of spacecraft

4 REVIEWS AND AUDITContractor shall have established review system/committees/boards in the organisation toensure review of designs, selection/procurement of parts & materials, manufacturingprocesses, test results/anomalies etc. Apart from regular progress/status meetings, few ofthe mandatory review are:

• Equipment qualification status review




• Preliminary/Baseline Design Review• Critical Design Review• Test readiness review• Test results review• Pre Delivery review• Flight model completion review including performance compliance verification• Contractor also shall have periodic & surprise audits of all the participating

fabrication & test units

5 EEE PARTS SELECTION & CONTROL• Only space qualified parts with heritage shall be used.• Only ESA/NASA/MIL certified parts shall be used• Establish a system to control selection, qualification, application and traceability of

parts• Parts older than 10 years shall not be used• New parts proposed shall be qualified & ISRO approval to be obtained before use.• Radiation hardness policy shall be spelt out. Minimum radiation design margin of

1.5 shall be ensured.• All parts shall have Radiation tolerance as below for GEO mission

– TID threshold of 100 K rads min.– LET threshold of 40 MeV-cm2/mg for SEU– LET threshold of 100 MeV-cm2/mg for SEL

• Parts sensitive for low dose rate shall be subjected to ELDRS (0.1 rad/min) test• Declared parts list shall be provided for all equipments

6 MATERIALS & PROCESS CONTROL• Only space qualified materials & processes with heritage shall be used.• Contractor shall establish a system to control selection, qualification, application

and traceability of materials & processes• New materials & processes proposed for use shall be qualified & ISRO approval to

be obtained before use.• All materials used shall have TML/RML of <1% & CVCM of <0.1%• Materials & processes used outside the spacecraft shall withstand high temperatures,

space radiation effects• Prohibited materials like cadmium, pure Tin, pure Zinc shall not be used.• Contractor shall generate approved/declared materials & processes list.

7 RELIABILITY ASSURANCETo ensure/demonstrate robustness of the design following reliability analyses shall be




carried out• Reliability apportionment/prediction• Parts stress derating analysis• Failure mode effects and criticality analysis• System level FTA• Thermal/mechanical analysis• Worst case circuit analysis• Risk analysis & impact analysis of residual risks• Wearout analysis• Radiation design margin analysis• Spacecraft charging analysis• SEU/SET analysis• Single point failure/critical item list and their control plan

8 SOFTWARE QUALITY ASSURANCEThe Product Assurance team shall be responsible for the implementation and control offlight software which includes.

• Software requirements review• Software design review• Code walk through• Software testing & embedded system level testing• Configuration & version control, traceability

All software used to produce & test flight hardware shall conform to plan

9 MODEL PHILOSOPHY AND TESTContractor shall establish/demonstrate equipment qualification status to ISRO, further testsare as follows

• Subsystem level– All new subsystems shall have DVM/QM and be qualified including life

demonstration test– Heritage subsystems with minor modifications-due to spec change, use

environment change- shall be subjected to protoflight level tests– Heritage subsystems with no changes shall be subjected to acceptance level

tests• System/spacecraft level

– Integrated spacecraft shall undergo protoflight level testsAll the test levels, test margins & test sequence will be mutually agreed as per ISROrequirements.

ISRO


ISRO Ka BAND HTS

SCHEDULE E

LAUNCH INTERFACES&

LAUNCH SUPPORT SERVICES

ISRO/ISAC/HTS/RFP/001/06VER-01 SCHEDULE - ELAUNCH INTERFACES & LAUNCH SUPPORT SERVICES

December 2015 Page E-1ISRO CONFIDENTIAL PROPRIETARY

TABLE OF CONTENTS

1 INTRODUCTION .............................................................................................................3

2 LAUNCH VEHICLE COMPATIBILITY ............................................................................3

2.1 CANDIDATE LAUNCH VEHICLES.....................................................................................3

3 LAUNCH SUPPORT SERVICES .......................................................................................4

3.1 SPACECRAFT PREPARATION FOR THE LAUNCH ..............................................................4

3.2 LAUNCH VEHICLE/SPACECRAFT INTERFACE VERIFICATION TESTING ...........................4

3.2.1 LAUNCH SITE FUNCTIONAL TEST REHEARSAL ..............................................................4

3.2.2 LAUNCH SITE FUNCTIONAL TEST ...................................................................................4

3.3 LAUNCH SITE SUPPORT...................................................................................................5

3.4 SPACECRAFT PROPELLANT AND PRESSURANT............................................................... 5

3.5 SAFETY ...........................................................................................................................5



1 INTRODUCTIONThis Schedule describes the Launch vehicle compatibility requirement and the launch supportservices that shall be provided by the Contractor in preparation for the launch of the Satellite.The Contractor will be required to perform all work necessary to ensure the compatibility ofthe Satellite design with the specified launch vehicle and shall provide the evidence necessaryto demonstrate and document such compatibility.

2 LAUNCH VEHICLE COMPATIBILITYThe Contractor shall be responsible for ensuring that the satellite is compliant with all theconstraints imposed by the launch vehicle (Operational, Safety, Environmental etc). TheCompatibility requirements refer to all interfaces and all other aspects concerned with theLaunch vehicles and launch sites including safety.

The Contractor shall also ensure that the Satellite is compatible with the mission profiles andthe mission sequences for the launch vehicles that will be defined by the launch vehicleagencies. The baseline reference for the compatibility requirements with the launch vehicleshall be the latest version of Launch Vehicle Users Manual.

Contractor shall carry out the Work necessary to ensure that the Spacecraft is compatible withthe Compatible Launch Vehicles as per section 2.1. This Work includes all the analyses, tests,logistics planning, licensing, manpower allocation and documentation necessary to complywith the Interface Requirements between Contractor and the Launch Vehicle Agency.Contractor shall support Customer and participate in the interface meetings with LaunchServices Contractor selected by Customer on an as-needed basis. The location of thesemeetings may vary, such as the Launch Services Contractor headquarters, Launch Site,Contractor's facilities or Customer' facilities.

Customer will specify for each launch vehicle integration meeting the disciplines for whichpresence of either the primary responsible or one of the deputies is required. Contractor shallsend personnel to the meetings according to this requirement.

2.1 CANDIDATE LAUNCH VEHICLESThe Contractor shall design the Spacecraft to be compatible with the following launchvehiclesa) Ariane-Vb) Protonc) Sea launchd) Atlase) Falcon 9 Heavy



3 LAUNCH SUPPORT SERVICES3.1 SPACECRAFT PREPARATION FOR THE LAUNCHContractor is responsible for shipping the Spacecraft, the associated launch preparation siteequipment, the Spacecraft propellants and pressurants, the staff required to perform the Workat the launch preparation site, the required Equipment to interface with the launch vehicle andall required GSE to the Spacecraft Launch Site. This excludes the Spacecraft Selected LaunchVehicle-to-Spacecraft adapter and separation system, which will be supplied by LaunchServices Contractor to Customer.

The shipment of the Spacecraft shall be done by the Contractor using appropriate means, withthe satellite fully integrated in a single container, with no need to disassemble or de-configurethe Spacecraft through the whole process of installing, transporting and removing theSpacecraft from its container.

Contractor is responsible for performing in conjunction with the Launch Services Contractorall the Work related to prepare the Spacecraft for launch, transportation, loading andunloading of spacecraft, and support the Launch Services Contractor with the data andcompatibility Work to allow the Launch Services Contractor to launch and to reachSuccessful Injection and Separation. Unloading of Spacecraft and GCS containers, as well astransportation devices used inside the airplane and inside the high bay (with exception ofcrane) are under responsibility of Contractor.

3.2 LAUNCH VEHICLE/SPACECRAFT INTERFACE VERIFICATION TESTING

Contractor shall perform comprehensive testing of the Spacecraft to verify compatibilitybetween the Spacecraft, the launch vehicle adapter and separation system, and the launchvehicle. The compatibility test shall include demonstration of compatibility with the launchvehicle and co-passenger (if allocable) electromagnetic compatibility requirements withadequate margins.

3.2.1 LAUNCH SITE FUNCTIONAL TEST REHEARSALContractor shall conduct a launch site functional test rehearsal and shall exercise allprocedures that will be executed at the Launch Preparation Site, utilizing the same testequipment that will be later on used to test the Spacecraft at the Launch Preparation Site.

3.2.2 LAUNCH SITE FUNCTIONAL TESTContractor shall perform a series of system Launch Site Functional Tests (LSFT) at the launchsite to verify that all Spacecraft subsystems are functionally unchanged as a result of shipmentand that the Spacecraft is fully functional and ready for launch. The LSFT shall constitute partof the pre-launch validations.



3.3 LAUNCH SITE SUPPORTContractor shall be responsible for the planning and implementing of all Spacecraft-relatedactivities at the Launch Site in accordance with the Launch Operations Plan, to be submittedby Contractor prior or at the first ground operations meeting, typically twelve (12) monthsprior to Delivery Date. The Launch Operations Plan consists, as a minimum, of a day-to-dayactivities schedule, manpower deployment schedule, a countdown command plan providinginformation on responsibilities and command chain during final countdown, communicationrequirements, battery management schedule, and a list of support equipment required byLaunch Services Contractor.

Contractor shall provide all necessary personnel, ground support and other equipment(including test equipment and spares), material and documentation needed at the Launch Site.

Contractor shall provide the capability to transmit the satellite telemetry to the MissionControl Center in real time during the final launch rehearsal and on the launch day.

3.4 SPACECRAFT PROPELLANT AND PRESSURANTThe required quantity of propellant typically MON3 + MMH and xenon in case of ElectricPropulsion System with sufficient spare will be provided by the Contractor as per the launchrequirements. The other fluid requirements need to be projected during the finalization oflauncher Procurement.

3.5 SAFETYThe Contractor shall demonstrate throughout the duration of the contract compliance with thesafety requirements of the agencies responsible for the designated launch vehicles by carryingout the required hazard analyses, safety analyses and submissions and test programmes.

ISRO


ISRO Ka BAND HTS

SCHEDULE F

MISSION AND POST LAUNCH SUPPORTSERVICES

ISRO/ISAC/HTS/RFP/001/06VER-01 SCHEDULE - F

MISSION AND POST LAUNCH SUPPORT SERVICES

December 2015 Page F-1ISRO CONFIDENTIAL PROPRIETARY

TABLE OF CONTENTSSCOPE ........................................................................................................................................3

1 MISSION SUPPORT SERVICES .......................................................................................3

1.1 GROUND CONTROL SYSTEM ...........................................................................................3

1.2 GROUND CONTROL SOFTWARE ......................................................................................4

1.3 DYNAMIC SATELLITE SIMULATOR (DSS) ........................................................................4

1.4 ORBITAL OPERATIONS (OOS) SOFTWARE .......................................................................5

1.5 GCS TEST REQUIREMENTS ..............................................................................................5

1.6 GROUND CONTROL SYSTEM PRODUCT ASSURANCE ......................................................5

2 TRAINING ......................................................................................................................5

3 POST LAUNCH SUPPORT SERVICES.............................................................................. 6

3.1 GEOSTATIONARY TRANSFER ORBIT ...............................................................................6

3.2 POST LAUNCH .................................................................................................................6

3.3 IN-ORBIT TEST ................................................................................................................6

3.4 OPERATINAL PHASE .......................................................................................................6




SCOPEContractor shall provide the support required to plan, prepare and carry out the operationsfunctions for the in-orbit control of the Satellite. This Schedule describes the pre-launchmission support activities including training, post launch activities, In-orbit testing andhandover operations.

1 MISSION SUPPORT SERVICES1.1 GROUND CONTROL SYSTEMThe Spacecraft will be controlled from satellite operations facilities in one of Customer’sTTC sites after handing over to Customer.

Contractor will be responsible for development and installation of the Real-time Software(RTS) used for the operational control of the Spacecraft after hand-over at the customercontrol center.

Contractor shall provide RTS in particular including the following aspects:a. Ground to Spacecraft Interface definitionb. TM/TC Database interpretation and conversionc. Spacecraft memory map and adaptation of the RTS on-board processor supportd. Spacecraft-specific Derived Parameterse. Interface of the Spacecraft Manufacturer’s Off-line Spacecraft Control Software to the

RTS (if applicable)f. Interface of the RTS (TM, Ranging and Spacecraft Procedures) with the Orbital

Operations software (Orbops) specificationsg. Procedures conversion into the RTS Commanding Languageh. Procedure validation on the RTS hardware

Contractor shall be responsible for the design, development, on-site integration, test anddocumentation of the following Contractor provided software and hardware components:

a. Any off-line Spacecraft-specific Software not explicitly mentioned in this Schedule,but required to properly operate the Spacecraft. This includes but not limited to anytools developed by the manufacturer which would aid in the on-orbit operation of thesatellite such as excel spreadsheet calculations and the like.

b. Dynamic Satellite Simulator (DSS) software;c. Validated TM/TC databased. Validated Spacecraft Procedurese. Specific Orbital Operations software (Orbops) specification.f. Transponder Management Softwareg. Set point determination tool




1.2 GROUND CONTROL SOFTWARE1. The Contractor furnished software shall perform its functions for the entire Spacecraft

lifetime, specifically during all Spacecraft modes including all orbit change maneuversand momentary switching transients.

2. The Man Machine Interface (MMI) shall be in English.

3. All system messages, failure and warning popup windows must be on top, i.e. not behidden by other windows.

4. The access control implemented in the delivered software components shall not rely on alist of hardcoded IP addresses.

Contractor shall be responsible for providing the following software packages:a. RTSb. Dynamic Satellite Simulator (DSS) for the Spacecraft;c. Orbital Operations software (Orbops) modules definitions for the Spacecraft;d. TM/TC database filese. Any deliverable software, with their source code when applicable, which is used to

properly operate any specific hardware delivered in the frame of this Agreement.

1.3 DYNAMIC SATELLITE SIMULATOR (DSS)This section defines the Technical and Operational Requirements for the Dynamic SatelliteSimulator (DSS) software package to be supplied to Customer. Contractor shall provide withthe DSS all the COTS licenses that are no longer commercially available at the time of S/CPDR + 3 months.

1. The DSS shall support connection and disconnection from the RTS without affecting therunning simulation

2. The DSS shall support a network outage and be able to reestablish a RTS reconnectiononce the network link is restored.

3. Connections attempted from an RTS session to a DSS already connected to a differentRTS session shall simply be rejected, and not affect the running simulation.

4. Attempts to start more than the allowed number of DSS sessions shall simply be rejected,and not affect the running simulations.

5. It shall be possible to start and operate the simulation from a remote computer.




1.4 ORBITAL OPERATIONS (OOS) SOFTWAREAll orbital control related operations shall be performed using the Contractor delivered orbitaloperations software (also referred to as “OOS software”). Consequently, this software shallbe upgraded as required to cover Spacecraft-specific operations which are currently notsupported.

1.5 GCS TEST REQUIREMENTS1. Testing of the delivered GCS software and hardware components shall consist of GCS

tests (Factory Acceptance Test), tests at Customer's premises (Site Acceptance Test, RTSCompatibility Test and End-to-End test) and at Contractor's premises (DSS FactoryAcceptance Test and Spacecraft Compatibility Test).

2. On-site acceptance testing will follow after the successful completion of the factoryacceptance test with the purpose of demonstrating that the GCS conforms to Customer’srequirements.

1.6 GROUND CONTROL SYSTEM PRODUCT ASSURANCEThe program shall utilize Contractor's established system of plans, procedures, and controlsthat are applied on ground station programs. These practices shall be applied consistentlywith a commercial level program and in accordance with ISO 9001.

2 TRAININGContractor shall submit for Customer’s approval a Training Plan covering the operations andsystem design of the Spacecraft. The Training Plan shall address the recommended timeframe for each of these courses within the program time schedule and outline the contents ofeach of the courses described below, including reference material, desirable traineebackground and any other pertinent information.

All courses shall be in English. Qualified engineers, analysts and programmers actuallyinvolved in the Spacecraft project shall give training to ensure a first-hand knowledgetransfer to Customer personnel.

Contractor shall test and evaluate the trainees during the training to assure adequate progress.Periodic reports of training progress shall be submitted to Customer.

Customer shall have video recording rights on all courses specified herein for its own futureinternal use.

A detailed training programme shall be submitted by contractor for review and acceptance bycustomer.




3 POST LAUNCH SUPPORT SERVICES3.1 GEOSTATIONARY TRANSFER ORBITContractor and Customer shall jointly define the final orbit injection parameters. To achievesuch objective, two Mission Design Reviews shall be held. In the Mission Design Reviews,Contractor and Customer shall jointly prepare the transfer orbit mission planning andanalysis.

3.2 POST LAUNCHFollowing Launch, Contractor shall provide all the necessary facilities, equipment andservices to acquire Spacecraft telemetry and orbit data, perform necessary calculations andanalyses, and required maneuvers to take the spacecraft from transfer orbit to final parkingslot. Customer representatives shall be part of the overall mission operations.

3.3 IN-ORBIT TESTContractor shall be responsible for the planning and execution of the bus and payload systemsin-orbit tests, to be executed from Customer’s satellite operations site. Contractor shall beresponsible for all interim station-keeping and station re-location maneuvers computationafter the transition from the transfer orbit mission to the nominal on-orbit mode. Contractor isresponsible for the entire in-orbit testing of the Spacecraft in accordance with the In-OrbitTest Plan submitted by Contractor for approval by Customer. Contractor will perform thecommunications payload in-orbit acceptance tests from Customer’s operations site.Contractor is responsible for the execution and analysis of the In-Orbit payload tests.

3.4 OPERATINAL PHASEAfter Launch and up until the end of the Spacecraft Propellant Life, Contractor shall providea continuous support to Customer to evaluate Customer-provided off-nominal data generatedfrom Spacecraft operations. Contractor shall respond by making as expeditiously as possiblea diagnostic of the provided data and indicate either the ability to continue to satisfy therequirements of this Agreement for the Design Life, or an inability or risk of an inability tosatisfy the requirements. In the latter case Contractor shall propose actions to minimize theimpact of the Spacecraft anomalous performance.

Contractor shall be responsible for reporting to Customer all anomalies that occurred in othersatellites manufactured by Contractor and any other deficiencies, like industry alerts, newlydiscovered fabrication or design deficiencies and all other anomalies not disclosed toCustomer before Launch of the Spacecraft that may affect the Spacecraft. Contractor isresponsible for disclosing such deficiencies within one week of their occurrence oruncovering by Contractor. Contractor is responsible for proposing to Customer actions toeliminate or ameliorate the deficiency, and maintain Spacecraft compliant to the requirementsof this Agreement throughout the Design Life.




Contractor shall deliver to Customer all updates and improvements on operationalprocedures, Spacecraft dynamic simulator and real-time and analysis operational tools,applicable to the Spacecraft and developed in the period between the launch of the Spacecraftand the end of Spacecraft Propellant Life. Dynamic Satellite Simulator updates shall belimited to the scope required to maintain compliance to the requirements of this Schedule.Contractor shall have previously validated these updates and improvements upon delivery toCustomer. Contractor shall support Customer in the validation of such updates andimprovements, and shall provide the reason for such updates and the proposed method toimplement and operate such updates and improvements.

ISRO


ISRO Ka BAND HTS

SCHEDULE G

COMMERCIAL & CONTRACTUALREQUIREMENTS

ISRO/ISAC/HTS/RFP/001/06VER-01 SCHEDULE - GCOMMERCIAL AND CONTRACTUAL REQUIREMENTS

December 2015 Page G-1

ISRO CONFIDENTIAL PROPRIETARY

TABLE OF CONTENTS1 INTRODUCTION............................................................................................................3

2 SCOPE OF WORK..........................................................................................................3

3 COMMERCIAL PROPOSAL........................................................................................3

4 PROPOSAL FORMAT & CONTENT..........................................................................4

4.1 TECHNO-COMMERCIAL BID .................................................................................................4

4.2 PRICE BID ............................................................................................................................5

5 PAYMENT TERMS / ADVANCE PAYMENT / BANK GUARANTEE...................6

6 TAXES AND DUTIES.....................................................................................................7

7 SECURITY DEPOSIT ....................................................................................................7

8 DELIVERY SCHEDULE ...............................................................................................7

9 REPLACEMENT SPACECRAFT.................................................................................7

10 INSURANCE....................................................................................................................8

11 LAUNCH SERVICES .....................................................................................................8

12 AWARD OF CONTRACT..............................................................................................8

13 WARRANTY: ..................................................................................................................9

14 PERFORMANCE BANK GUARANTEE.....................................................................9

15 TERMS OF CANCELLATION .....................................................................................9

16 LIQUIDATED DAMAGES ............................................................................................9

17 FORCE MAJEURE: .....................................................................................................10

18 RETENTION OF PROPOSALS..................................................................................10

19 CONFIDENTIALITY ...................................................................................................11

20 APPROVALS, GOVERNMENT & REGULATORY CLEARANCES...................11

21 INTEGRITY...................................................................................................................11

22 AMENDMENT OF RFP DOCUMENT ......................................................................11

23 RIGHTS ON THE SPACECRAFT..............................................................................12

24 ARBITRATION.............................................................................................................12

ANNEXURE-1 : PRICE BID................................................................................................15

ANNEXURE-2 : MAJOR MILESTONES FOR REALIZATION OF SPACECRAFT.16




ANNEXURE-3 : ADVANCE PAYMENT BANK GUARANTEE....................................17

ANNEXURE-4 : BANK GUARANTEE FOR SECURITY DEPOSIT ............................20

ANNEXURE-5 : PERFORMANCE BANK GUARANTEE..............................................23




1 INTRODUCTIONContractors are invited to submit a Firm Fixed Price Proposal (referred to as FFP) for twosatellites called ISRO Ka HTS 1 and ISRO Ka HTS 2, which is specified in Schedule A to Gof this RFP.

2 SCOPE OF WORKThe broad scope of work includes

Development, design, construction, integration, testing of the satellites. Necessary launch compatibility studies and supporting for launcher co-ordination

with the selected launcher agency Shipment to the launch preparation site Testing at the launch preparation site, preparations for launch including fueling, all

work required in conjunction with the launch services contractor for launch vehiclecompatibility

Satellite installation on the launcher and launch support Conduction of low earth orbit operations In orbit testing of the satellites Hardware and software for ground segment validation and satellite operations Storage of the satellites, if required.

3 COMMERCIAL PROPOSALThe proposal should be submitted in two separate sealed covers containing:

Part-1 : Techno-commercial bidPart-2 : Price bid

The techno-commercial bid shall also include a format of the price bid (masking the pricevalues) indicating the various options and description of line items being offered by thecontractor.

The price bid shall give a detailed break-up including details for both ISRO Ka HTS-1 andISRO Ka HTS-2.

The Validity of proposal and quotation should be nine (9) months from the date ofopening the tender.




4 PROPOSAL FORMAT & CONTENT4.1 TECHNO-COMMERCIAL BIDTechno-commercial bid (Masking the price values) shall be given in both hard and soft formi.e. in CD media only with one backup (Note: Submission in no other electronic format shallbe considered)

Any electronic media (CD) sent from abroad through courier or post needs to be cleared byIndian Customs authorities, which might result in the delayed delivery of consignment. Inorder to avoid the delay, the contractor is advised to send the hard copies and CDs in separateconsignments.

The proposal shall consist of the following parts:

a. Volume I. Executive Summary Summary of satellite design Company experience – Historical Delivery Performance in similar satellite should be

highlighted. Comparison of satellite main parameters (power, mass, etc) with other satellites of

the same product line This volume shall not contain any commercial information

b. Volume II. Response to RFP SchedulesThe detailed and elaborated responses to the schedules A to G of this RFP may beprojected.

c. Volume III. Compliance statementThe contractor shall provide an Item wise Compliance for the schedule A to G of thisRFP. In case of partial compliance of any specification / schedule, the contractor shallpropose reason / justification / alternate-solution indicating why the particular partialcompliance will not impact the overall performance of the satellite. Customer’s decisionin this regard will be final.

d. Volume IV. Commercial ProposalThe contractor shall provide the price bid in the format enclosed in Annexure-1 withmasking of the price information.

Part-1: Techno-Commercial should not contain any price details and be kept in a sealed coverduly super scribing as under:




QUOTATION AGAINST RFP NO. - ISRO/ISAC/HTS/RFP/001/06

DUE ON ………………………………..at 1400 Hrs IST

PART I - TECHNO-COMMERCIAL BID for ISRO Ka HTS

SPACECRAFT

The cover should indicate “SENDER” address.

4.2 PRICE BIDPrice bid only one copy in hard form.

The contractor shall provide the price bid in the format enclosed in Annexure-1 of thisschedule. This part shall indicate the Price in FE (Preferably in US Dollar) (with detailedbreak-up of applicable taxes and duties if applicable).

The detailed break-up for each item of supply or services shall be provided. Wheneveroptions are quoted, the same should also be indicated with quantity and unit rate separately.The prices are to be mentioned both in figures and in words. In the event of any discrepancyin the price between figures and words, the price indicated in words shall only prevail andvalid.

The Price offered shall be binding for a period of nine (9) months from the tender openingdue date.

This part should also be kept in a sealed cover super scribing as follows:

QUOTATION AGAINST RFP NO. - ISRO/ISAC/HTS/RFP/001/06

DUE ON ………………………………..at 1400 Hrs IST

PART II - PRICE BID for ISRO Ka HTS SPACECRAFT

The Two sealed covers prepared as above should be kept in another envelope, sealed andsuper scribed as under:-




QUOTATION AGAINST RFP NO. – ISRO/ISAC/HTS/RFP/001/06

DUE ON ………………………………..at 1400 Hrs IST

for ISRO Ka-HTS SPACECRAFT

and addressed to:

Head, Purchase and StoresISRO Satellite Centre, HAL Airport RoadBangalore – 560017,Karnataka, IndiaPh No. : +918025084004/+918025084369Fax No. : +918025205283/+918025205284E-mail : [email protected]

Please note that proposals submitted after the above deadline will not be considered in ourevaluation and will eliminate the contractor from the competition out of fairness to thecontractors who delivered their proposals on time.

5 PAYMENT TERMS / ADVANCE PAYMENT / BANK GUARANTEEThe Contractor shall clearly bring out the payment terms in their proposal. Advance payment,required if any, shall be against the milestone events of the contractual progress and shouldbe safeguarded with an equivalent amount of bank guarantee. The contractor shall identifysuch milestones while proposing the advance payment terms. Typical milestones are listed inAnnexure-2 for reference.

In case of advance payment requested, the bids are evaluated by suitably loading interest @11% per annum on the advance amount (or) applicable rate declared by the Government ofIndia which are in force at the time of advancing.

The Bank Guarantee should be valid from the date of advance / milestone payment up toTHREE (3) MONTHS after the completion of Spacecraft Flight Module Completion Review(FMCR). The cost of the Bank Guarantee shall be borne by the selected contractor.

Format of Advance Payment Bank Guarantee is as per Annexure 3.




6 TAXES AND DUTIES(a) Applicable taxes and duties may be specified in the bid. Applicable Income tax at source(TDS) with or without DTAA, wherever applicable shall be effected from payments of thesuccessful contractor.(b) Customs Duty (Applicable for HTS-2): ISRO, Satellite Centre, Bengaluru is exemptedfrom payment of Customs Duty vide Notification No.12/12-Cus dated 17.03.2012. Thenecessary Customs Duty Exemption Certification [CDEC] shall be provided. Tenderer[s] arerequested to take note of this aspect and submit the quotation clearly mentioning that thequoted price does not include Customs Duty and ISAC has to provide CDEC.(c) All costs pertaining to custom clearance and other incidentals shall be borne by thecontractor while transporting subsystems, components etc to India for HTS-2. Customer willhowever provide end user certificate and all possible document support in obtaining customclearance.

7 SECURITY DEPOSITOn acceptance of the contract, the Contractor shall, at the option of the Customer and withinthe period specified by him, deposit with him, in cash or in any other form as the Customermay determine, security deposit not exceeding ten percent of the value of the Contract. TheSecurity Deposit shall be obtained through Bank Guarantee or Fixed Deposit Receipt fromany of Scheduled Bank, executed on non-judicial stamp paper of appropriate value and shallbe valid for a period of 60 Days beyond the date of completion of the Contract (FMCR). Ifthe Contractor is called upon by the Customer to deposit, ‘Security’ and the Contractor failsto provide the security within the period specified, such failure shall constitute a breach of theContract, and the Customer shall be entitled to make other arrangements for the re-purchaseof the Contracted Satellite at the risk of the Contractor and/or to recover from the Contractor,damages arising from such cancellation. Format of Security Deposit Bank Guarantee is as perAnnexure - 4.

8 DELIVERY SCHEDULEThe ISRO Ka HTS-1spacecraft is required to be ready for shipment to launch base within orlatest by 36 months from the date of signing the contract. ISRO Ka HTS-2 Spacecraft shall beready 6 months after the FMCR of ISRO Ka HTS-1spacecraft.

9 REPLACEMENT SPACECRAFTIn case of launch failure and the option of replacement spacecraft being exercised byCustomer, Contractor shall make all efforts to deliver the spacecraft at the earliest.




10 INSURANCEThe contractor will be responsible for the insurance of the spacecraft till delivering thesatellite at the launch site. Customer will be responsible for Launch All Risk (LAR) insuranceand post-launch in-orbit insurance.

11 LAUNCH SERVICESCustomer will be responsible for the procurement of Launch services. The contractor shallprovide satellite compatible to the Launch Vehicles as proposed in Schedule E, Section 2.1.

12 AWARD OF CONTRACTThe evaluation criteria will include, but not limited to, technical compliance, payment termsand price. Adequate preference will be given to the contractors who have in flight experienceand heritage of similar High Power / High Throughput spacecraft. The contractor shouldaddress all these aspects in the proposal. Participation of Customer engineers is also theessence. Item 4.3 of Schedule-A shall be followed for more detail in this regard. The travel,living and transportation expenses for ISRO engineers will be borne by the customer.

Customer reserves all the rights to award a Contract or Contracts for the whole or part of thework required by the RFP or to make no award at all.

The selection of winning bid is at the sole discretion of Customer.

The cost of preparing proposals and participating in the whole process in response to thisRFP shall be borne solely by the Contractors. The release of this RFP does not create anyfinancial obligation on the part of Customer.

Customer shall assign the overall responsibility of implementation on a single contractor(prime contractor). Any dependency on any sub-contractors shall be managed by the primecontractor and should not have any bearing whatsoever on Customer and the performance ofthe final contract.

The offer should also contain the details of the sub-contract, if any, proposed to be awardedby the Contractor for some part of the system or subsystem to another supplier/Contractor.The details should include information like work/business profile of such a supplier,experience in executing/supplying similar type of system/subsystem for which thesubcontract is being awarded, etc.




13 WARRANTY:(a) The product offered shall be warranted by the Contractor for a nominal period of 18months from the date of acceptance of the product by the Customer (FMCR).(b) The term Warranty period, shall nominally mean the period of 18 months from the dateon which the Spacecraft or allied materials has been accepted by the Customer (FMCR).

In case the launch of the satellite is delayed, the warranty period with mutual agreement willbe accordingly extended.

14 PERFORMANCE BANK GUARANTEEThe Contractor shall, furnish a Bank Guarantee (prescribed format as per Annexure-5) from aNationalized/Scheduled Bank approved by the Customer for an amount equivalent to 10% ofthe value of the Contract after completion of the FMCR for the due performance of thesatellite in the warranty period. On completion of the warranty period, the Performance BankGuarantee will be returned to the Contractor without any interest.

15 TERMS OF CANCELLATIONIn the normal circumstances, customer does not foresee short closure / cancellation of thecontract either in part or full. However, under extreme circumstances, customer shall reservethe right to short close / cancel the contract by giving a prior notice of 60 days without havingany financial implications on either side of the parties to the concluded contract.

16 LIQUIDATED DAMAGESIf the Contractor fails to deliver the Stores (Satellite and associated deliverables in this RFP)within the time specified in the Contract or any extension thereof, the Purchaser shall recoverfrom the Contractor as Liquidated Damages a sum one-half of one percent (0.5 percent) ofthe Contract price of the undelivered Stores for each calendar week of delay or part thereof.The total Liquidated Damages shall not exceed Ten percent (10 percent) of the Contract priceof the unit or units so delayed. Stores will be deemed to have been delivered only when allits component/parts/deliverables are also delivered in full to make use of the wholesystem/equipment as the case may be. If certain components/parts are not delivered in time,the entire value of Contract/Stores will be considered as delayed until such time as themissing parts are delivered.




In case of delay in delivery of the stores beyond the delivery date stipulated in the PurchaseOrder/Contract or any extension thereof, such stores shall be received under protest. Thestores shall be accepted without prejudice to the terms and conditions of the Contract.

17 FORCE MAJEURE:(a) Neither party shall bear responsibility complete or partial non performance of any of hisobligations [except for failure to pay any sum which has become due on account of receipt ofgoods under the provisions of the purchase order/contract] if the non-performance resultsfrom such force majeure circumstances such as, but not restricted to, flood, fire, earthquake,civil commotion, sabotage, explosion, epidemic, quarantine restriction, strike, lock out,freight embargo, hostility, acts of public enemy and other acts of God as well as war orrevolution, military operation, blockade, or any other circumstances beyond the control of theparties that have arisen after the conclusion of purchase order/contract. In suchcircumstances, the time stipulated for the performance of an obligation under the purchaseorder/contract may be proportionately extended.

(b)The party for whom it has become impossible to meet the obligation under this contractdue to force majeure condition, will notify the other party in writing not later than 21 daysfrom the date of commencement of unforeseeable event. Unless agreed by both the parties,in writing, the contractor shall continue to perform his obligations under the purchaseorder/contract as far as is practical and shall seek all reasonable alternative means forperformance not prevented by the force majeure event.

(c)Any Certificate issued by the Chamber of Commerce or any other competent authority ororganization of the respective country shall be sufficient proof of commencement andcessation of the above circumstances. In case of failure to carryout complete or partialperformance of an obligation for more than 60 days, either party shall reserve the right toterminate the contract totally or partially. A prior written notice of 30 days to the other partywill be given informing of the intention to terminate any liability.

18 RETENTION OF PROPOSALSAll documents submitted in response to the RFP shall become the whole & sole property ofcustomer. Any information in such documents that is proprietary to the Contractor should bespecified clearly. At its own discretion, customer reserves the right to disclose proposals, ifrequired, including proprietary information, to appropriate personnel / consultants selected bycustomer to assist it in the evaluation.




19 CONFIDENTIALITYAny technical information passed on to the contractor shall be treated as confidential by theContractor and shall not be disclosed directly or indirectly to third parties, without the priorconsent of customer.

20 APPROVALS, GOVERNMENT & REGULATORY CLEARANCESIt shall be the responsibility of the contractor to obtain all the necessary clearances from itsgovernment for submission of technical data related to the proposal and for supplying theSpacecraft and in-orbit technical support as and when required. No act of the Government inits sovereign capacity shall be considered as an event of Force Majeure in the contract to befinalized. The Contractor shall submit a declaration of acceptance in this regard along withthe bid and shall obtain the necessary governmental clearances before signing of the contractin the event of being chosen as the successful Contractor.

Any other contractual / statutory requirements either governmental or procedural issueswhich influence the contract implementation by the customer may be explicitly brought out.The contractor shall obtain all the necessary governmental licenses that are required to beprovided for supply and maintenance of stated Spacecraft.

21 INTEGRITYNo company or person (other than a full-time bona-fide employee working solely for theContractor), including any local coordinator, shall be employed or retained by the Contractorto solicit, secure or participate in any contract resulting from this RFP or shall participate inany negotiations that may be called for.The Contractor shall not pay to any company or individual, any fee, commission orbrokerage, contingent upon or resulting from the award of any contracts resulting from thisRFP. The Contractor shall explicitly make a statement to this effect in its proposal.

22 AMENDMENT OF RFP DOCUMENTAt any time prior to the deadline for submission of Applications, customer either on its ownor on request of the contractors may amend the RFP Document by issuing addenda. To givethe contractors reasonable time to take an addendum into account in preparing theirresponses, customer may, at their discretion, extend the deadline for the submission ofproposals.




23 RIGHTS ON THE SPACECRAFTThe contractor at no point can claim any right on the spacecraft. Customer is the sole ownerof the spacecraft. The contractor should operate the spacecraft based on the instructionsprovided by ISRO.

24 ARBITRATIONAny dispute of difference whatsoever arising between the parties out of or relating to theconstruction, meaning, scope, operation or effect of this contract or the validity or the breachthereof shall be settled by arbitration in accordance with the Rules of Arbitration of theIndian Council of Arbitration and the award made in pursuance thereof shall be binding onthe parties.

Further terms and conditions and detailed mechanisms on arbitration shall be worked outjointly by ISRO and the selected bidder at the time of finalizing the contract.




Annexure




Annexure-1

PRICE BID

This part shall indicate the Price in FE (Preferably in US Dollar) (with detailed break-up ofapplicable taxes and duties).

Contractor shall propose a detailed break-up considering all the aspects of spacecraftrealization against the typical format given below. This can be considered as a baselineformat.

S.N. Item

Price in FE(Preferably in US$)

HTS-1 HTS-2

1. Development, design, construction of the satellite2. AIT activities3. Shipment to the Launch base and Pre-Launch

activities4. Launch Early Operation Phase (LEOP)5. In-Orbit Testing (IOT)6. Ground segment hardware and software support

requirements7. Spacecraft Storage, if required8. Any other Item

Total




Annexure-2

MAJOR MILESTONES FOR REALIZATION OFSPACECRAFT

S.N. Milestone TimeSchedule

Payment ExpectedHTS-1 HTS-2

1 Signing of the contract T02 Spacecraft EQSR3 Preliminary Design Review (PDR)4 Critical Design Review (CDR)5 Thermovac completion6 Spacecraft Dynamic test and CATF

completion7 Flight Module Completion Review

(FMCR) and authorization to shipmentT0+36Months

HTS-1 FMCR+ 6 Months

8 Spacecraft preparation in Launch base9 Launch10 IOT completion11 Completion of one year In-orbit operation




Annexure-3

ADVANCE PAYMENT BANK GUARANTEE

Bank Guarantee No: Date:

THIS GUARANTEE AGREEMENT is made at on thisday of in favor of the President of India

represented by Director, ISRO Satellite Centre, Department of Space, Government of India,(hereinafter referred to as "The Purchaser").

WHEREAS, on the day ofM/s. having its registered office at

(hereinafter referred to as"The Contractor") entered into a contract bearing No. (the Contract) withthe President of India acting through the Director, ISRO Satellite Centre, Department ofSpace, Government of India, HAL Airport Road, Vimanapura P.O., Bangalore – 560 017(hereinafter referred to as "The Contractor") for sale of the

& related items, services & activities for a total sum ofUSD (US Dollars )

AND WHEREAS, under the terms and conditions of the contract, payments are to be madeby the Purchaser to the contractor as specified in the contract.

AND WHEREAS, the Purchaser has agreed in pursuance of the said terms and conditions ofthe contract to make _ % payment as advance to the Contractor on their furnishing aBank Guarantee in the manner herein contained.

We (the Bank) Bank, in consideration of thePurchaser having agreed to pay to the Contractor advance payment of US Dollars herebyagree unequivocally and unconditionally to pay within 48 (forty eight) hours without demurany amount up and not exceeding 100% of the payment made to the Contractor irrespectiveof existence of “Force Majeure” conditions and acts of government including denial ofExport License or Subsequent sanctions after receipt of Export License, upon Purchaser’sgiving notice of the encashment demand in writing by its authorized officer, which noticeshall include the certification that a) Contractor has defaulted under the Contract failed to




fulfill its obligation to deliver goods in accordance with the terms of the Contract referred toin the Advance Payment Bank Guarantee and has declined the purchaser’s written request to(a) extend the effective period of this Guarantee for an additional term of One HundredTwenty (120) Days or until delivery, acceptance and shipment of Goods whichever is earlier,or (b) arrange for the issuance of a replacement Advance Payment Bank guarantee with aterm of One Hundred Twenty (120) Days from expiration of the existing Advance PaymentGuarantee or until delivery acceptance, and shipment of Goods whichever is earlier, withoutit being necessary for the Purchaser to adduce any proof in support of the claim.

AND WE (the Bank) hereby agree not to revoke thisGuarantee during its currency except with the previous consent of the Purchaser in writingand agree that any change in the constitution of the said Contractor or the Guarantor onForce Majeure clauses of Contract and acts of Government shall not discharge the Guarantorfrom its liability hereunder.

AND WE (the Bank)_ hereby agree that our (Bank's)liability hereunder shall not be discharged by virtue of any arrangement between thePurchaser and the Contractor whether with or without knowledge/or consent, or by reason ofthe Purchaser showing any indulgence or forbearance to the Contractor, whether as topayment, time performance or any other matter whatsoever which, but for this provision,would amount to discharge of the surety under any law.

THE GUARANTEE herein contained shall remain in full force and effect from the date of itsexecution in the first instance until the earlier of (a) final delivery, shipment, acceptance ofproducts or (b) T0 (defined as effective date of contract) + months and orContractor’s presentment to Bank of Purchaser’s Certificate of acceptance of

.

Unless a claim under the Guarantee is presented to us (Bank) prior to date of expiration ofthis Guarantee, all rights of the Purchaser under the guarantee shall be forfeited and we [TheBank] shall be relieved and discharged from all liabilities hereunder.

The present Guarantee will, therefore lapse on the earlier of (a) the final delivery, acceptanceand shipment of the product or (b) T0 (defined as effective date of Contract)months. However, in the event that one of the contracting parties instituted ArbitrationProceedings with the ICADR, India, and that the latter informs us thereof in writing duringthe period of validity of the presenting Guarantee, then, the validity of the present guaranteewill be automatically extended until the end of the calendar month which follows the monthduring which the final decision of the ICADR is taken.




In witness whereof (Bank)2016.

have executed this, the day of

Bank




Annexure-4

BANK GUARANTEE FOR SECURITY DEPOSIT

(From Indian Agents on behalf of foreign principals)

(On non-judicial stamp paper of appropriate value)

1. WHEREAS on or about the .................... (date), M/S. .........................................having its Office at ................................. (hereinafter referred to as “the Contractor”)entered into a Contract bearing No. .........................dated ....................... (hereinafterreferred to as “The Contract”) with the President of India (hereinafter referred to as “theGovernment”) for supply of ...................................... (hereinafter referred to as “theEquipment”). The Contract recognizes that M/S. ................................... as the IndianAgent of the Contractor in India who will furnish the Security Deposit bond on behalf ofthe Contractor for satisfactory performance of the Contract as per the terms andconditions contained in the said Contract.

2. AND WHEREAS under the terms and conditions of the Contract, an amount of`.............................. (Rupees ................................... only) representing around .............%of the Contract value is to be furnished as Bank Guarantee by the Indian Agent on behalfof the Contractor, in a manner herein contained, duly executed by the................................. (Bank) towards satisfactory performance of the Contract.

3. We, ................................................ (name of the Bank & address) [hereinafter referredto as “the Bank'] at the request of M/S. ......................................, authorised Indian Agentof the Contractor, do hereby undertake to pay to the Accounts Officer of the……………………………………(name & address of the Centre/Unit) on behalf of theGovernment (hereby referred to as ‘the Accounts Officer’), an amount not exceeding`................... (Rupees .................................. only) against any loss or damage caused to orsuffered or would be caused to or suffered by the Government by reason of any breach bythe said Contractor of any of the terms and conditions contained in the said Contract.

4. WE, ...................................... (Bank) do hereby undertake to pay the amount due andpayable under this guarantee without any demur, merely on a demand from the saidAccounts Officer, stating that the amount claimed is due by way of loss or damage causedto or would be caused to or suffered by the Government by reason of breach by the said




Contractor of any of the terms and conditions contained in the said Contract or by reasonof the Contractor’s failure to perform the said Contract. Any such demand made on theBank shall be conclusive as regards the amount due and payable by the Bank under thisguarantee. However, our liability under this guarantee shall be restricted to an amount notexceeding `.............. (Rupees ....................................... only).

5. WE, .................................................. (Bank) undertake to pay to the Government anymoney so demanded notwithstanding any dispute or disputes raised by the Contractor, inany suit or proceedings pending before any Court or Tribunal relating thereto, ourliability under this present guarantee being absolute and unequivocal. The payment somade by us under this bond shall be valid discharge of our liability for paymentthereunder and the Contractor shall have no claim against us for making such payment.

6. WE, ................................ (Bank) further agree that the guarantee herein containedshall remain in full force and effect during the period that would be taken for theperformance of the said Contract and that it shall continue to be enforceable till all thedues of the Government under or by virtue of the said Contract have been fully paid bythe said Contractor and its claims satisfied or discharged or till the said Accounts Officercertifies that the terms and conditions of the said Contract have been fully and properlycarried out by the said Contractor and accordingly discharges this guarantee. Unless ademand or claim under this guarantee is made on us in writing on or before........................ (date) [two months beyond the Contract completion date], we shall bedischarged of all liabilities under this guarantee thereafter.

7. WE, .................................... (Bank) further agree that the Government shall have thefullest liberty without our consent and without affecting in any manner our obligationshereunder to vary any of the terms and conditions of the said Contract or to extend timefor performance by the said Contractor from time to time or to postpone for any time orfrom time to time any of the powers exercisable by the Government against the saidContractor and to forbear or enforce any of the terms and conditions relating to the saidContract and we shall not be relieved from our liability by reason of any such variation/sor extension being granted to the said Contractor or for any forbearance, act orcommission on the part of the Government or any indulgence by the Government to thesaid Contractor or by any such matter or thing whatsoever which under the law relating tosureties would, but for this provision, have effect of so relieving us.

8. THIS GUARANTEE will not be discharged due to the change in the constitution ofthe Bank or the Contractor.

9. WE, ................................... (Bank) lastly undertake not to revoke this guaranteeduring its currency except with the previous consent of the Government in writing.




Place :

Date :

…………………………..………………………….

(Signature of the Authorised Officer of the Bank)

…………………………………………

…………………………………………

(Name and designation of the officer)

Seal

Name, Address of the Bank (Head Office) with Phone/Fax Nos.

Name & Address of the Branch with Phone/Fax Nos.




Annexure-5

PERFORMANCE BANK GUARANTEE

1. WHEREAS on or about the ................... day of ................. (month & year),M/S. .......................... having its Registered Office at ............................ (hereinaftercalled 'The Contractor') entered into an Agreement No. .................... dated .................(hereinafter called 'The Contract') for manufacture and supply of ......................... with thePresident of India (hereinafter called 'The Purchaser').

2. AND WHEREAS, under the terms and conditions of the Contract final paymentamounting to US$.............. (US$ ..................... only) under the Contract is to be madeagainst a performance bond in the form of bank guarantee furnished by the Contactor fora sum of US$............. (US$ ........................ only) equivalent to 10% (ten per cent) of thevalue of the Contract towards satisfactory performance of the ....................... (hereinaftercalled 'the Equipment') valid for a period of 12 months from the date of putting intooperation of the said equipment or ............. months from the date of receipt of the last lotof consignment whichever is earlier (specify as per warranty clause in the PurchaseOrder).

3. NOW WE, ............................... (name & address of the Bank) in consideration of thepromises and payment of the final/balance amount of US$............. (US$ ................ only)under the Contract to the Contractor hereby agree and undertake to pay on demand andwithout any demur to the Accounts Officer, ............................ (name of the Centre/Unit)on behalf of the Purchaser (hereinafter referred to as the Accounts Officer), a sum notexceeding US$............... (US$ ...................... only) against any loss or damage that maybe suffered by the Purchaser by the reasons of any unsatisfactory performance of the saidequipment.

4. AND WE, ...................... (Bank) hereby also agree that the decision of the saidAccounts Officer as to whether the said equipment is giving satisfactory performance ornot and as to the amount of loss or damages suffered by the Purchaser on account ofunsatisfactory performance of the said equipment shall be final and binding on us.




5. AND WE, ......................... (Bank) hereby further agree that our liability hereundershall not be discharged by virtue of any agreement between the Purchaser & theContractor whether with or without our knowledge and/or consent or by reason of thePurchaser showing any indulgence or forbearance to the Contractor whether as topayment, time for performance or any other matter whatsoever relating to the Contractwhich, but for this provision, would amount to discharge of the surety under the law.

6. OUR GUARANTEE shall remain in force until .................. (two months beyondContract warranty period) and unless a claim under the guarantee is lodged with us on orbefore the above date, all rights of the Purchaser under the Guarantee shall be forfeitedand we shall be relieved and discharged from all our liabilities hereunder. Our liabilityunder this guarantee shall not be affected by any change in our constitution or theconstitution of the Contractor.

Dated the .............. day of ............................ (month & year).

For ....................................

(Name of the Bank)

Postal address of the Bank and

Fax No.& e-Mail id)

Seal of the Bank

ISRO

REQUEST FOR PROPOSAL (RFP)

FOR

ISRO Ka BAND HTS

ANNEXURE-A

COMPLIANCE MATRIX

ISRO/ISAC/HTS/RFP/001/06VER-01 COMPLIANCE MATRIX

December 2015 CM-1ISRO CONFIDENTIAL PROPRIETARY

TABLE OF CONTENTS

1 SCHEDULE A: STATEMENT OF WORK ....................................................................................................................................................3

2 SCHEDULE B: TECHNICAL SPECIFICATION REQUIREMENTS ....................................................................................................... 6

3 SCHEDULE C: ASSEMBLY INTEGRATION TESTING AND SATELLITE STORAGE PLAN .......................................................15

4 SCHEDULE D: PRODUCT ASSURANCE REQUIREMENTS................................................................................................................. 16

5 SCHEDULE E: LAUNCH INTERFACES AND LAUNCH SUPPORT SERVICES ...............................................................................16

6 SCHEDULE F: MISSION AND POST LAUNCH SUPPORT SERVICES............................................................................................... 17

7 SCHEDULE G: COMMERCIAL AND CONTRACTUAL REQUIREMENTS.......................................................................................18



COMPLIANCE MATRIXNote to contractor: Compliance for tables and figures wherever applicable under given clause should also be stated by the contractor.

1 SCHEDULE A: STATEMENT OF WORK

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NotApplicable Remarks

Section I1 Introduction2 Scope

Section II1 Introduction2 Flight Satellites2.1 Definitions2.1.1 Spacecraft2.1.2 Compatible Launch Vehicle2.1.3 Operational Life2.1.4 Design Life2.1.5 Propellant Life2.1.6 Beginning of Life (BOL)2.1.7 End of Life (EOL)3 Responsibility3.1 ISRO Ka HTS-13.2 ISRO Ka HTS-23.3 ISRO Band HTS-1 and HTS-23.4 Risk Manageent




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3.5 Anomalies Handling & Documentation3.6 Export Authorizations

4 Realization Plan & Customer'sTechnical Visibility and Access

4.1 Realization Plan4.2 Timeline and Procurement4.3 ISRO Participation4.4 Heritage4.5 Access to Workplace4.6 Spacecraft Life and Mass4.6.1 Spacecraft Life4.6.2 Spacecraft Mass

5Hardware and Software for GroundSystem Validation Testing and SatelliteOperations

5.1 TCR Suitcase5.2 Telemetry Data5.3 Satellite Simulator5.4 Payload Reconfiguration Software Tool5.5 Steerable Antenna Control Software Tool5.6 ADCS Re-Programming Tool5.7 Training

5.8 Warranty and Maintenance PostAcceptance

5.9 Tele-command Encryption6 Launch Vehicle Compatibility7 Ground Control Compatibility




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8 Product assurance Requirements9 Satellite Reviews and Schedule

9.1 Equipment qualification Status Review(EQSR)

9.2 Design Reviews9.2.1 Preliminary Design Reviews(PDR)9.2.2 Critical Design Review (CDR)

9.3 Manufacturing, Assembly, Integration andTest Reviews

9.3.1Satellite Qualification Status Review(SQSR)/ Environmental Test ReadinessReview (ETRR)

9.3.2 Flight Model Completion Review (FMCR)9.3.3 Test Reviews9.3.4 In Orbit Acceptance Review (IOAR)10 Non Deliverable Equipment11 Spare Provisions12 Satellite Storage and Safe Keeping13 Services13.1 Launch Support Services13.2 Mission Support services

Section III1 Program Management2 Major Realization Milestones



2 SCHEDULE B: TECHNICAL SPECIFICATION REQUIREMENTS


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1 Communication Payload Requirements1.1 Introduction1.1.1 Spectrum Allocation1.1.1.1 User Spectrum1.1.1.2 Hub Spectrum1.1.2 Coverage1.1.3 Beam layout

1.1.3.1 User Beams Configurationa) Frequency Reuse Scheme

1.1.3.2 Hub Beams Configuration1.1.4 Payload Configuration1.1.4.1 KA-Band Beacon1.1.4.2 Antenna RF Tracking System

1.1.4.3Payload Redundancy

a) Ka-Band Transpondersb) Ka-Band Beacon Transmitters

1.1.5 Payload Performance Specifications1.1.5.1 Effective Isotropic Radiated Power (EIRP)1.1.5.2 G/T1.1.5.3 Bandwidth1.1.5.4 Polarization1.1.5.5 Multi Beam System C/I1.1.5.6 Coverage Requirement1.1.5.7 Transponder Modes of operation




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a) Fixed Gain Mode (FGM)b) Automatic Level Control (ALC)c) Mute Mode

1.1.5.8 Saturation Flux Density Range1.1.5.9 ALC Mode O/P Power Variation1.1.5.10 Noise Power Ratio at Given OBO

1.1.5.11 EIRP Stability (Excluding AntennaEffects)

1.1.5.12 G/T Stability (Excluding Antenna Effects)1.1.5.13 EIRP Stability (Including Antenna Effects)

1.1.5.14 Small Signal Gain Stability (ExcludingAntenna Effects)

1.1.5.15 G/T Stability (Including Antenna Effects)

1.1.5.16 Multi-beam System C/I (Over 100%Coverage Area)

1.1.5.17 Wide Band Receive Response Rejection1.1.5.18 Commandable Gain in FGM1.1.5.19 ALC Mode Specifications1.1.5.20 Frequency Conversion1.1.5.21 Linearity of Receive Section1.1.5.22 In-Band Frequency Response (Saturation)

1.1.5.23 In-Band Frequency Response ( SmallSignal)

1.1.5.24 Input Gain Slope Response1.1.5.25 Total Gain Slope Response1.1.5.26 In-Band Group Delay (Input)1.1.5.27 In-Band Group delay (Total)




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1.1.5.28 Out of Band Receive NarrowbandRejections

1.1.5.29 Out of Band Transmit NarrowbandRejections

1.1.5.30 Total Channel Phase-Shift

1.1.5.31 AM/PM Transfer & ConversionCoefficient

1.1.5.32

Spurious Output1) Spurious Within Transmit Band2) Spurious Outside Transmit Band3) Spurious due to Receive LO

1.1.5.33 Overdrive1.1.5.34 Switching Interrupts1.1.5.35 TWTA Cathode Life1.1.5.36 Unit Interdependence

1.1.5.37

High Power Considerationsa) Multipactionb) Passive Inter-Modulation Productsc) EMI/EMC

1.1.5.38 Cessation of Emissions1.1.5.39 Parameters for Throughput Estimation1.1.6 Ka Band Beacon Specifications1.1.6.1 Transmit Frequency1.1.6.2 EIRP1.1.6.3 Coverage1.1.6.4 Polarization1.1.6.5 Cross Polarization Discrimination (XPD)




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1.1.6.6 Frequency Stability1.1.6.7 EIRP Stability1.1.7 Specifications of IOT-Horn

1.1.8 Optional Configuration (India Beams +Steerable Beams)

1.1.8.1 Payload Performance Specifications

2 Telemetry, Command andRanging(TCR) System

2.1 General Requirements2.2 Command Subsystem2.2.1 General Requirements2.2.2 Configuration2.2.3 Frequency of operation2.2.4 Frequency Stability2.2.5 Antenna Coverage2.2.6 Antenna Polarization2.2.7 Antenna Axial Ratio2.2.8 Input Power Flux Density2.2.9 Modulation2.2.10 Baseband Requirement2.2.11 Command Encryption (Optional)2.3 Telemetry Subsystem2.3.1 General Requirements2.3.2 Configuration2.3.3 Frequency of Operation2.3.4 Frequency Stability2.3.5 Antenna Coverage




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2.3.6 Antenna Polarization2.3.7 Antenna Axial Ratio2.3.8 Effective Isotropic Radiated Power (EIRP)2.3.9 EIRP Stability2.3.10 Modulation2.3.11 Base band Requirements2.4 Ranging Subsystem2.4.1 General Requirements2.4.2 Range Tone Frequencies and Modulation2.4.3 Range Calibration2.4.4 Modulation2.4.5 Ranging Performance Requirement

2.5 Ground Station Compatibility Unit(Suitcase Model)

3 Electrical Power System3.1 General Requirements3.2 Solar Array3.3 Batteries3.4 Power Control & Protections3.5 Redundancy3.6 Power Budget4 Thermal Requirements4.1 Thermal Range and Temperatures4.2 Thermal Design

4.3 Thermal Mathematical Models andAssociated Documentation

4.4 Thermal Subsystem Development Testing




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4.4.1 Analysis/Measurement

4.4.2

Heat Pipe/Loop heat Pipe DevelopmentTests

a) Unit levelb) Heat pipe/Loop heat pipe assembly

4.4.3Deployed Thermal Radiator

a) Unit levelb) Heat pipe/Loop heat pipe assembly

4.4.4 Spacecraft Thermal Balance Test4.5 Thermal Control Material And Processes5 Structure Subsystem5.1 Functional Requirements5.2 Performance Requirements5.3 Design Requirements5.3.1 Environmental Loads5.3.2 Design Criteria

6 Attitude Determination and ControlSystem (ADCS)

6.1 Definition of Spacecraft Axes

6.2 Attitude Control System FunctionalRequirements

6.3 Attitude Determination6.4 Attitude Control6.4.1 Stability6.4.2 Pointing Accuracies6.4.3 Station Keeping Maneuvers6.4.4 In-Orbit Antenna Pattern Measurement




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6.4.5 Control Bias Capability6.4.5.1 Operation in Inclined Orbit at the EOL6.4.5.2 Ground Control6.4.5.3 Unit and Reference Switching6.4.6 Safe Modes and Re-acquisition6.4.7 Payload Shedding in Emergency6.4.8 Mass Properties6.4.9 Autonomy6.4.10 Control Electronics Fault protection6.4.11 Propulsion Interfaces6.4.12 Telemetry Interfaces6.4.13 Specific ADCS Test6.4.14 Satellite On Board Processor Memory Data6.4.15 ADCS Block Diagrams6.4.16 Attitude Control and Determination Data6.4.17 Fuel Sloshing Data7 Propulsion System7.1 Subsystem Requirements7.2 Operational Requirements7.2.1 Reliability7.2.2 Safety7.2.3 Operational Lifetime7.2.4 Storage Life7.2.5 Maintainability7.2.6 Ground Operations7.2.7 Launch




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7.2.8 Transportability7.2.9 Filling and Draining7.3 Design and Construction7.3.1 Parts, Materials and Processes7.3.1.1 Lubricants7.3.1.2 Dissimilar Metals7.3.1.3 Fungus Resistance7.3.1.4 Compatibility With Corrosive Materials7.3.2 Mechanical Design7.3.3 Thermal Design Description8 Mechanisms and Pyrotechnics

8.1

Design Requirementsa) Rotating Devices Requirementsb) Deployment and Pointing

Mechanism Requirements8.2 Pyrotechnic Devices9 Orbital operations9.1 Transfer Orbit9.1.1 Attitude Error Budget9.1.2 Tracking Network9.2 Station Keeping9.2.1 Terminology9.2.2 Requirements9.2.3 Drift and Eccentricity Control Maneuvers9.2.4 Inclination Control Maneuvers

9.2.5 Hybrid Inclination/Eccentricity ControlManeuvers




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9.3 Additional Requirements9.3.1 Spacecraft Longitude Relocation9.3.2 Ranging System Requirements

9.3.3 Maximum Delta -V for Attitude RecoveryManeuvers

9.3.4 Disposal Orbit and Passivation9.4 Propellant Budget10 Design and Analyses10.1 Design Analyses and Study Reports10.1.1 Requirements10.1.2 Analyses Data10.2 System Level Analyses10.2.1 Dynamic Analyses10.2.2 Propellant Budget Analyses10.2.3 Mission Analyses10.2.4 Mass Properties Analyses

10.2.5 Electromagnetic Compatibility (EMC)Analyses

10.2.6 Environmental Effects Analyses10.2.6.1 Radiation Effects Analyses10.2.6.2 Satellite Charging Analyses10.2.6.3 Satellite Contamination Analyses10.2.6.4 Satellite Depressurisation Analyses10.2.7 Payload Analyses10.2.7.1 Communication Link Analyses10.2.7.2 Antenna Pointing Error Analyses10.2.8 FDIR Response and Observability




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10.3 Subsystem Level Analyses

10.3.1 Telemetry, Tele-command and RangingAnalyses

10.3.2 Thermal Analyses10.3.3 Structural Analyses10.3.4 Electrical Power Analyses

10.3.5 Attitude Determination and ControlAnalyses

10.3.6 Propulsion Analyses

3 SCHEDULE C: ASSEMBLY INTEGRATION TESTING AND SATELLITE STORAGE PLAN


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1 Subsystem Level Testing2 System Level Integration Tests2.1 AIT Sequence2.2 Ground Support Equipments2.3 Additional Test Requirements2.3.1 Electromagnetic Compatibility test2.3.2 ESD Tests2.3.3 Alignments2.3.4 Interface Verification Tests

2.3.5 Spacecraft RF, Base Band and SoftwareCompatibility Testing

3 Spacecraft Hardware Transportationand Handling

3.1 Safety of all Operations




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3.2 Safety of Mechanical Operations3.3 Safety of Electrical Operations3.4 Safety of Thermal Vacuum Operations4 Spacecraft Storage Plan

4 SCHEDULE D: PRODUCT ASSURANCE REQUIREMENTS


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1 Program Content2 Product Assurance Requirements

3 Spacecraft Product AssuranceManagement

4 Reviews and Audit5 EEE Parts Selection & Control6 Materials & Process Control7 Reliability Assurance8 Software Quality Assurance9 Model Philosophy and Test

5 SCHEDULE E: LAUNCH INTERFACES AND LAUNCH SUPPORT SERVICES


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1 Introduction2 Launch Vehicle Compatibility




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2.1 Candidate Launch Vehicles3 Launch Support Services3.1 Spacecraft Preparation for the Launch

3.2 Launch Vehicle/Spacecraft InterfaceVerification Testing

3.2.1 Launch Site Functional Test Rehearsal3.2.2 Launch Site Functional Test3.3 Launch Site Support3.4 Spacecraft Propellant and Pressurant3.5 Safety

6 SCHEDULE F: MISSION AND POST LAUNCH SUPPORT SERVICES


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1 Mission Support Services1.1 Ground Control System1.2 Ground Control Software1.3 Dynamic Satellite Simulator (DSS)1.4 Orbital Operations (OOS) Software1.5 GCS Test Requirements1.6 Ground Control System Product Assurance2 Training3 Post Launch Support Services3.1 Geostationary Transfer Orbit3.2 Post Launch




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3.3 In-Orbit Test3.4 Operational Phase

7 SCHEDULE G: COMMERCIAL AND CONTRACTUAL REQUIREMENTS


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1 Introduction2 Scope of Work3 Commercial Proposal4 Proposal Formal & Content

4.1

Techno-Commercial Bida) Volume I. Executive Summaryb) Volume II. Response to RFP

Schedulesc) Volume III. Compliance Statementd) Volume IV. Commercial Proposal

4.2 Price Bid

5 Payment Terms/Advance Payment/BankGuarantee

6 Taxes and Duties7 Security Deposit8 Delivery Schedule9 Replacement Spacecraft10 Insurance11 Launch Services




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12 Award of Contract13 Warranty14 Performance Bank Guarantee15 Terms of Cancellation16 Liquidated Damages17 Force Majeure18 Retention of Proposals19 Confidentiality

20 Approvals, Government & RegulatoryClearances

21 Integrity22 Amendment of RFP Document23 Rights on the Spacecraft24 ArbitrationAnnexure - 1 Price Bid

Annexure - 2 Major Milestones for Realization ofSpacecraft

Annexure - 3 Advance Payment Bank GuaranteeAnnexure - 4 Bank Guarantee for Security DepositAnnexure - 5 Performance Bank Guarantee

Ab-1

ANNEXURE-B

ABBREVIATIONS

ADCS Attitude Determination and Control SubsystemAGC Automatic Gain ControlAIT Assembly Integration and TestingALC Automatic Level ControlAM/PM Amplitude Modulation to Phase ModulationAMF Apogee Motor FiringAMPL Approved Materials and Processes ListAOCS Attitude Orbit Control SystemAPSK Amplitude & Phase Shift KeyingAPL Authorized Parts ListAQL Acceptable Quality LevelASIC Application Specific Integrated CircuitBDR Baseline Design ReviewBER Bit Error Rate BOL

Beginning of LifeBPE Beam Pointing ErrorC/I Carrier to Interference RatioCAB Corrective Action BoardCATF Compact Antenna Test FacilityCCSDS Consultative Committee for Space Data SystemCDMO Configuration Data Management officeCDR Critical Design ReviewCIDL Configuration Identification Data ListCIDL Configuration Identification Document ListCQSR Component Qualification Status ReviewCRB Change Review BoardCSI Customer Source InspectionCW Continuous WaveDB DecibelDCL Declared Component ListDOD Depth of DischargeDoCon Down Converter

Ab-2

DRB Delivery Review BoardDSS Dynamic Satellite SimulatorDVB Digital Video BroadcastDVM Design Verification MatrixDVB-S2 Digital Video Broadcasting – Satellite Version 2EDA Effective Date of AgreementEDC Effective Date of ContractEED Electro Explosive DevicesEEE Electrical, Electronic & ElectromechanicalEGSE Electrical Ground Support EquipmentEIDP End Item Data packageEIRP Effective Isotropic Radiated PowerEMC Electromagnetic CompatibilityEMI Electromagnetic InterferenceEOC End of ChargeEOL End of LifeEPC Electronic Power ConditionerEQSR Equipment Qualification Status ReviewESD Electrostatic DischargeETRR Environmental Test Readiness ReviewFDIR Fault Detection Identification and ReconfigurationFDMA Frequency Division Multiple AccessFGM Fixed Gain ModeFMCR Flight Model Completion ReviewFMECA Failure Mode Effect Criticality AnalysisFOV Field of ViewFPGA Field Programmable Gate ArrayFR Flight ReadinessFRB Failure Review BoardFRR Flight Readiness ReviewFSK Frequency Shift KeyingFTA Fault Tolerance AnalysisG/T Antenna Gain to Noise TemperatureGCS Ground Control SystemGIDEP Government-Industry Data Exchange ProgramGMT Greenwich Mean TimeGSE Ground Support EquipmentGSID Ground to Spacecraft Interface DocumentGSO Geostationary Satellite Orbit

Ab-3

GTD Geometric Theory of DiffractionGTO Geostationary Transfer OrbitHDL High Definition LanguageHPA High Power AmplifierHRF Harmonic Reject FilterHSIA Hardware Software Interaction AnalysisHTS High Throughput SatelliteIBO Input Back-offICD Interface Control DocumentIMP Intermodulation ProductIOAR In-Orbit Acceptance ReviewIOT In-Orbit TestIST Integrated Satellite TestITU International Telecommunication UnionLAM Liquid Apogee MotorLAT Lot Acceptance TestLCAMP Linearized Channel AmplifierLCTWTA Linearized Channel Amplifier (LCAMP) & TWTALEOP Launch and Early Operation PhaseLH Linear HorizontalLHCP Left-Handed Circular PolarizationLNA Low Noise AmplifierLO Local OscillatorLSFT Launch Site Functional TestsLV Linear VerticalLVT Lot Validation TestMF-TDMA Multi Frequency – Time Division Multiple AccessMGSE Mechanical Ground Support EquipmentMIC Microwave Integrated CircuitMIP Mandatory Inspection PointsMMI Man Machine InterfaceMMIC Monolithic Microwave Integrated CircuitModCod Modulation & CodingMRB Material Review BoardMRR Mission Readiness ReviewMUX MultiplexerNCR No Calibration RequiredNPR Noise Power RatioNTP Normal Temperature and Pressure

Ab-4

O/P OutputOBO Output Back-offOOS Orbital Operations softwareOPCA Output Power Control AttenuatorORBOPS Orbital Operations softwarePA Product AssurancePAD Parts Approval DocumentPAP Product Assurance ProgramPDR Preliminary Design ReviewPFM Proto Flight ModelPIMP Passive Intermodulation ProductPMP Parts, Materials and ProcessPMPCB Parts, Materials, and Process Control BoardPPAE Program Product Assurance EngineersPPAM Program Product Assurance ManagerPQSR Product Qualification Status ReviewPSF Pre-Select Filter PSK

Phase Shift KeyingPWB Printed Wiring BoardsQAM Quadrature Amplitude ModulationQCI QUALITY Conformance InspectionQML Qualified Material ListR&QA Reliability and Quality AssuranceRADLAT Radiation Lot Acceptance testRCB Radiation Control BoardRF Radio FrequencyRFP Request for ProposalRFW Request for WaiverRHCP Right-Handed Circular PolarizationRMS Root Mean SquareRTOS Real Time Operating SystemRTE Real Time ExecutiveRTS Real time SoftwareRVT Radiation Verification TestSADA Solar Array Drive AssemblySCDRL Subcontractor Data Requirements ListSCTV Spacecraft Level Thermo VacuumSEE Single Event Effects

Ab-5

SET Single Event TransientSEU Single Event UpsetSFD Saturation Flux DensitySFT System Functional TestSGRP Single Ground Reference PointSMR Software Modification RequestsSOC State of ChargeSPF Single Point FailuresSPL Summary Parts ListSPR Software Problem ReportsSQSR Satellite Qualification Status ReviewSRI Standard Repair InstructionSRR Systems Requirements ReviewSSB Single Side BandSSPA Solid State Power AmplifierSTA Static Timing AnalysisSWM Switch MatrixTBD To be DecidedTC TelecommandTCP Transfer Control ProtocolTCR Telemetry, Command and RangingTCVs Technological Characterization VehicleTDMA Time Division Multiple AccessTID Total Ionization DoseTM TelemetryTMM Thermal Mathematical ModelTRB Test Review BoardsTRR Test Readiness ReviewsTVAC Thermo-VacuumTWT Travelling Wave TubeTWTA Travelling Wave Tube AmplifierULPC Uplink Power ControlVSWR Voltage Standing Wave RatioWCA Worst Case Analysis

REQUEST FOR PROPOSAL · 3.1 isro ka hts-1 For HTS-1 the contractor is fully responsible for all...

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Transcript of REQUEST FOR PROPOSAL · 3.1 isro ka hts-1 For HTS-1 the contractor is fully responsible for all...