taurus launch system september 1999 - brown …...release 3.0 september 1999 i thank you for...

September 1999

Release 3.0

Taurus® Launch SystemPayload User's Guide

Copyright © 1996—1999 by Orbital SciencesCorporation. All Rights Reserved.

®

Release 3.0 September 1999 i

Thank you for considering Orbital’s Taurus launch system to fulfill your launch service needs. This revision to the Taurus User's Guide is designed to familiarize mission planners and designers with the current capabilities and interfaces of the Taurus launch system, with the goal of enabling potential customers to perform mission feasibility trade studies and complete preliminary mission designs.

Orbital’s Taurus launch system has been designed to provide safe, successful, on-time launch services for a wide variety of payloads and mission designs. Our completely successful flights (March of 1994, and February and October of 1998) validated the Taurus system and operational concept as well as provided a solid foundation for our on-going Taurus Program. The Taurus team is in place, is fully dedicated to delivering a high quality launch service and is committed to mission success at all levels. In addition, Orbital’s Launch Systems Group has achieved and is operating under ISO-9000 certification.

Orbital designed the Taurus launch service to satisfy the varied requirements of our payloads, since the Taurus vehicle is simply a tool to enable the payload to accomplish its mission. Orbital offers a range of payload accommodations and interfaces to provide flexibility in designing a particular mission. In addition, Taurus ground and launch operations have been structured to streamline vehicle integration efforts. Originally designed for rapid response from austere launch sites for Department of Defense missions, the aspects of the Taurus launch system that enable these rapid operations have distinct advantages for commercial and scientific missions as well.

Orbital’s Taurus program team is eager to assist in your mission design and mission planning activities. This User’s Guide is intended to provide the basic information to allow potential customers to perform preliminary assessments of Taurus' suitability for their mission. Please contact the Taurus Program Office at any stage of your mission design process for further information and a mission-specific assessment.

Thank you in advance for considering Taurus for your launch service needs.

William A. WrobelTaurus Program ManagerOrbital/Launch Systems Group

Taurus User's Guide Preface

Release 3.0 September 1999 ii

This Taurus® Launch System Payload User's Guide is intended to familiarize potential customerswith the Taurus launch system, its capabilities and its associated services. The launch servicesdescribed herein are available for commercial procurement directly from Orbital SciencesCorporation. This document is not intended to be an entirely comprehensive description of allaspects of the Taurus launch service. Readers desiring further information on Taurus or on themethods of procuring Taurus launch services should direct their inquiries to:

Orbital Launch Systems Business DevelopmentOrbital Sciences Corporation

21700 Atlantic BoulevardDulles, Virginia 20166

Telephone: (703) 406-5227 Facsimile: (703) 406-3505

This document will be revised periodically. Recipients are encouraged to e-mail us at [email protected] or call us at our New Business number, (703) 404-7400, to ensure they willbe included on the mailing list for future updates.

Release 3.0 September 1999

TABLE OF CONTENTS

SECTION PAGE

1.0 INTRODUCTION .....................................................................................................1-12.0 DESCRIPTION OF TAURUS LAUNCH SERVICES............................................................. 2-1

2.1 Taurus Launch System Overview ............................................................................2-12.2 Taurus Launch Service ............................................................................................2-32.3 Description of the Taurus Vehicle ...........................................................................2-32.4 Primary Launch Support Equipment ........................................................................2-6

2.4.1 Launch Support Van ......................................................................................2-72.4.2 Launch Equipment Van .................................................................................2-8

3.0 GENERAL PERFORMANCE ...............................................................................................3-13.1 Vehicle Performance ...............................................................................................3-13.2 Mission Description ................................................................................................3-13.3 Taurus Ascent Profile ..............................................................................................3-1

3.3.1 Direct Ascent ...............................................................................................3-23.3.2 Parking Orbit Ascent ...................................................................................3-3

3.4 LEO Performance Capability ...................................................................................3-33.5 Injection Accuracy ................................................................................................3-113.6 Payload Deployment Attitude Options ..................................................................3-113.7 Separation Tip-Off Rates and Velocities ................................................................3-11

4.0 PAYLOAD ENVIRONMENTS ............................................................................................4-14.1 Payload Environments .............................................................................................4-14.2 Steady State and Transient Accelerations ................................................................4-14.3 Sine Vibration .........................................................................................................4-24.4 Natural Frequency Requirements ............................................................................4-34.5 Random Vibration ...................................................................................................4-34.6 Acoustic Environment .............................................................................................4-44.7 Pyro Shock ..............................................................................................................4-44.8 Thermal and Humidity ............................................................................................4-44.9 Contamination Control ............................................................................................4-54.10 RF Environment .....................................................................................................4-6

5.0 PAYLOAD INTERFACES ....................................................................................................5-15.1 Payload Interfaces ...................................................................................................5-15.2 Payload Fairings .....................................................................................................5-1

5.2.1 Payload Dynamic Envelope .........................................................................5-15.2.2 Payload Access Door...................................................................................5-1

5.3 Payload Mechanical Interface .................................................................................5-15.4 Electrical Interfaces .................................................................................................5-5

5.4.1 Payload Separation Connectors ...................................................................5-55.4.2 Umbilical Harness Pass-Throughs ...............................................................5-55.4.3 Payload Command and Control ..................................................................5-5

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TABLE OF CONTENTS(CONTINUED)

SECTION PAGE

5.4.4 Payload Discrete Status Monitoring .............................................................5-65.4.5 Pyrotechnic Initiation Signals ......................................................................5-75.4.6 Payload Analog Telemetry Monitoring ........................................................5-75.4.7 Serial Communication Interface ..................................................................5-75.4.8 Payload Battery Charging ............................................................................5-75.4.9 Pre-Launch Electrical Constraints ................................................................5-75.4.10 Pre-Separation Electrical Constraints ...........................................................5-75.4.11 Electrical Launch Support Equipment ..........................................................5-8

5.5 Payload Design Constraints .....................................................................................5-85.5.1 Payload Center of Mass Constraints .............................................................5-85.5.2 Final Mass Properties Accuracy ...................................................................5-95.5.3 Payload EMI/EMC Constraints .....................................................................5-95.5.4 Payload Dynamic Frequencies ....................................................................5-95.5.5 Payload-Supplied Separation Systems..........................................................5-95.5.6 System Safety Constraints ............................................................................5-9

6.0 GROUND AND LAUNCH OPERATIONS .........................................................................6-16.1 Taurus/Payload Integration Overview......................................................................6-16.2 Facilities .....................................................................................................6-1

6.2.1 Western Range Facilities .............................................................................6-16.2.1.1 Payload Processing .......................................................................6-1

6.2.1.1.1 Security ........................................................................6-36.2.1.1.2 Thermal and Cleanliness ..............................................6-36.2.1.1.3 Cranes ..........................................................................6-36.2.1.1.4 Electrical .......................................................................6-36.2.1.1.5 Communications ..........................................................6-46.2.1.1.6 Payload Fueling Capability ...........................................6-46.2.1.1.7 Other Equipment and Services ...................................... 6-4

6.2.1.2 Stage Integration Facility ...............................................................6-46.2.1.3 Launch Site Facilities ....................................................................6-4

6.2.1.3.1 Launch Support Equipment ...........................................6-56.2.1.3.2 Electrical .......................................................................6-66.2.1.3.3 Communications ..........................................................6-66.2.1.3.4 Security ........................................................................6-6

6.2.2 Eastern Range Facilities ...............................................................................6-76.3 Vehicle and Payload Integration Operations ...........................................................6-7

6.3.1 Launch Vehicle Integration ..........................................................................6-76.3.2 Payload Processing Operations ...................................................................6-86.3.3 Payload Transport to the Launch Site .........................................................6-11

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SECTION PAGE

6.4 Launch Operations ................................................................................................6-126.4.1 Facility and LSE Checkout .........................................................................6-126.4.2 Delivery of Taurus to the Launch Site ........................................................6-126.4.3 Payload Arrival and Activities at the Launch Site .......................................6-136.4.4 Integrated Testing and Vehicle Mate..........................................................6-146.4.5 Vehicle Arming .........................................................................................6-146.4.6 Launch ...................................................................................................6-15

6.4.6.1 Launch Control Organization .....................................................6-156.4.6.2 Rehearsals...................................................................................6-17

7.0 MISSION INTEGRATION ..................................................................................................7-17.1 Management Approach ...........................................................................................7-17.2 Orbital Mission Responsibilities ..............................................................................7-1

7.2.1 Customer Mission Responsibilities ...............................................................7-37.3 Working Groups .....................................................................................................7-37.4 Mission Planning and Development ........................................................................7-47.5 Reviews .....................................................................................................7-47.6 Documentation .....................................................................................................7-47.7 Customer-Provided Documentation ........................................................................7-6

7.7.1 Payload Questionnaire ................................................................................7-67.7.2 Payload Finite Element Model .....................................................................7-67.7.3 Safety Documentation .................................................................................7-67.7.4 PPF Requirements ........................................................................................7-67.7.5 Program Requirements Document (PRD) Mission Specific

Annex Inputs ...............................................................................................7-67.7.6 Payload Vibroacoustic Model ......................................................................7-77.7.7 Payload Thermal Model for Integrated Thermal Analysis .............................7-77.7.8 Test Verified Payload Finite Element Model ................................................7-77.7.9 Launch Operations Requirements (OR) Inputs .............................................7-77.7.10 Payload Launch Site Test Procedures ...........................................................7-77.7.11 Payload Integrated Test Procedure Inputs ....................................................7-77.7.12 Payload Final Mass Properties .....................................................................7-77.7.13 Taurus/Payload ICD Verification Documentation ........................................7-7

7.8 Orbital Produced Documentation ...........................................................................7-87.8.1 Payload Separation Analysis ........................................................................7-87.8.2 PPF Requirements Document ......................................................................7-87.8.3 Program Introduction (PI) ............................................................................7-87.8.4 Coupled Loads Analysis (Preliminary and Final) ..........................................7-87.8.5 Preliminary Mission Analysis .......................................................................7-9

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SECTION PAGE

7.8.6 Interface Control Document (ICD) ...............................................................7-97.8.6.1 Payload Mechanical Interface Control Drawing (MICD) ............... 7-97.8.6.2 Payload Electrical Interface Control Drawing (EICD) ....................7-9

7.8.7 RF Compatibility Analysis ............................................................................7-97.8.8 Vibroacoustic Analysis ................................................................................7-97.8.9 Integrated Missile System Pre-Launch Safety Package (MSPSP) ....................7-97.8.10 PRD Mission Specific Annex Inputs ...........................................................7-107.8.11 Integrated Thermal Analysis ......................................................................7-107.8.12 Fairing Clearance Analysis ........................................................................7-107.8.13 OR Inputs ..................................................................................................7-107.8.14 Integrated Launch Site Test Procedures......................................................7-107.8.15 Mission Constraints Document ..................................................................7-107.8.16 Final Countdown Procedure ......................................................................7-107.8.17 Final Mission Analysis ...............................................................................7-107.8.18 Taurus/Spacecraft ICD Verification Documentation ..................................7-107.8.19 Post-Launch Analyses ................................................................................7-11

8.1 OPTIONAL SERVICES .....................................................................................................8-18.2 Dual Manifest Options ............................................................................................8-1

8.2.1 63" Dual Payload Attach Fitting ...................................................................8-18.2.2 50" Aft Payload Capsule ..............................................................................8-1

8.3 Payload Fairing Options ..........................................................................................8-28.3.1 Additional Fairing Access Doors ..................................................................8-28.3.2 Fairing Access Doors of Non-Standard Size .................................................8-28.3.3 Fairing Acoustic Blanket Tailoring ...............................................................8-2

8.4 Optional Mechanical Interfaces ..............................................................................8-38.4.1 High Capacity Payload Separation System...................................................8-38.4.2 23" Payload Separation System....................................................................8-38.4.3 Non-Separating Payload Interface ...............................................................8-38.4.4 Separation System Thermal Treatment .........................................................8-3

8.5 Payload Isolation System .........................................................................................8-48.6 Performance Enhancements ....................................................................................8-48.7 Contamination Control Options ..............................................................................8-4

8.7.1 Class 10,000 Integration Environment .........................................................8-48.7.2 Fairing Surface Cleanliness Options ............................................................8-48.7.3 Instrument Purge or Spot Cooling System ....................................................8-48.7.4 Hydrocarbon Monitoring .............................................................................8-7

8.8 Real-Time Telemetry Downlink Through Payload Separation.................................. 8-7

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SECTION PAGE

8.9 Payload Propellant Loading .....................................................................................8-78.10 On-Board Video System ..........................................................................................8-7

APPENDIX A TAURUS PAYLOAD QUESTIONNAIRE ..................................................... A-1

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LIST OF FIGURES

FIGURE PAGE

2-1 Taurus Vehicle Configuration ........................................................................................2-12-2 Taurus Flight History and Current Manifest ...................................................................2-22-3 Transportable Launch Support Equipment .....................................................................2-22-4 Overview of Taurus Launch Service ..............................................................................2-32-5 Taurus Expanded View..................................................................................................2-42-6 Taurus Configuration Numbering Convention ...............................................................2-52-7 Taurus Propulsion Data Table .......................................................................................2-52-8 Taurus LSE Configuration ..............................................................................................2-62-9 LSV Layout ....................................................................................................................2-72-10 LEV Layout ....................................................................................................................2-83-1 Typical Taurus Direct Ascent Mission Profile ................................................................3-23-2 Typical Taurus Parking Orbit Mission Profile ................................................................3-33-3 Taurus Performance to 28.5° Orbits - CCAS Launch .....................................................3-43-4 Taurus Performance to 38° Orbits - CCAS Launch ........................................................3-53-5 Taurus Performance to 60° Orbits - South VAFB Launch...............................................3-63-6 Taurus Performance to 70° Orbits - South VAFB Launch...............................................3-73-7 Taurus Performance to 90° Orbits - VAFB Launch ........................................................3-83-8 Taurus Performance to Sun Synchronous Orbits - VAFB Launch ...................................3-93-9 Launch Range and Site Selection Are Tailored to Mission Performance

and Safety Requirements .........................................................................................3-103-10 Typical Stage Vacuum Impact Points for Taurus Launches from VAFB and CCAS .......3-103-11 Injection Accuracies to Low Earth Orbits (3-σ) ............................................................3-114-1 Phasing of Dynamic Loading Events ..............................................................................4-14-2 Payload Design C.G. Limit Load Factors .......................................................................4-24-3 Range of Sine Vibration Levels ......................................................................................4-24-4 Range of Shock Response Spectrums ............................................................................4-34-5 Payload Interface Random Vibration Levels ..................................................................4-44-6 63" Diameter Fairing Maximum Flight Level Payload Acoustic Environment ................4-54-7 92" Diameter Fairing Maximum Flight Level Payload Acoustic Environment ................4-64-8 Maximum Flight Level Pyroshock at the Payload Interface ............................................4-54-9 Taurus RF Sources .........................................................................................................4-65-1 Payload Dynamic Envelope for 63" Diameter Fairing ....................................................5-25-2 Payload Dynamic Envelope for 92" Diameter Fairing ....................................................5-35-3 Taurus Separation System..............................................................................................5-45-4 Interface Flanges for Taurus' Standard Separation Systems ............................................5-45-5 Typical Layout of Taurus Separation System (38.81" System Shown with Forward Ring

partially Removed to Show Lower Ring) ....................................................................5-45-6 38.81"/37.15" Separation System Structural Capability ..................................................5-55-7 Taurus Electrical Interface Block Diagram .....................................................................5-6

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LIST OF FIGURES(CONTINUED)

FIGURE PAGE

5-8 Payload Cone Structural Capability ...............................................................................5-85-9 Maximum Payload Induced Flight Pyroshock Levels at the Payload Interface ............... 5-96-1 Astrotech VAFB Payload Processing Facility..................................................................6-26-2 Astrotech VAFB PPF Details ..........................................................................................6-36-3 Astrotech PPF Power Sources ........................................................................................6-46-4 Missile Assembly Building.............................................................................................6-56-5 SLC-576: Primary Launch Facility for Taurus at the Western Range .............................. 6-66-6 LC-46 Launch Complex ................................................................................................6-76-7 LC-46 Tower ................................................................................................................. 6-76-8 Hardware Flow - Factory to Launch Site .......................................................................6-96-9 Vehicle Processing Flow at the MAB ...........................................................................6-106-10 Stage Integration in the MAB .......................................................................................6-116-11 Payload and LV Processing Flow at the PPF ................................................................6-116-12 Fairing Stowed in PPF .................................................................................................6-126-13 Payload Mated to Payload Cone .................................................................................6-126-14 First Fairing Half Installation .......................................................................................6-136-15 Payload Encapsulation Complete ................................................................................6-136-16 ECE Transport to Launch Site.......................................................................................6-136-17 Launch Site Processing Flow .......................................................................................6-146-18 ECE in Position for Integration With the Launch Vehicle .............................................6-146-19 Upper Stack Lift ..........................................................................................................6-156-20 Taurus Ready for Launch.............................................................................................6-166-21 Final Countdown Sequence ........................................................................................6-166-22 Typical Launch Decision Flow ....................................................................................6-176-23 Typical VAFB RCC Seating Arrangement .....................................................................6-187-1 Taurus Program Organization .......................................................................................7-17-2 Orbital Mission Responsibilities ....................................................................................7-27-3 Mission Integration Management Structure ...................................................................7-37-4 Customer Mission Responsibilities ................................................................................7-37-5 Typical Mission Schedule..............................................................................................7-57-6 Customer-Provided Documentation ..............................................................................7-67-7 Orbital Produced Documentation .................................................................................7-88-1 92" Fairing with 63" DPAF ............................................................................................8-18-2 63" Fairing with APC .....................................................................................................8-28-3 High Capacity Payload Separation System Structural Capability ...................................8-38-4 23.25" Payload Separation System Structural Capability ...............................................8-48-5 Payload Dynamic Envelope for 63" Diameter Fairing, STAR 37FM Configuration ......... 8-58-6 Payload Dynamic Envelope for 92" Diameter Fairing, STAR 37FM Configuration ......... 8-68-7 Taurus Performance to High Energy Trajectories ...........................................................8-7

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Release 3.0 September 1999 x

Taurus User's Guide Acronyms List

AIT Assembly and Integration TrailerAPC Aft Payload CapsuleCCAS Cape Canaveral Air StationCDR Critical Design ReviewCCTV Closed Circuit TelevisionCG Center of GravityCVCM Collected Volatile Condensable

MaterialDADS Dynamic Analysis and Design Sys-

temDARPA Defense Advanced Research

Projects AgencydB DecibelDC Direct CurrentDPAF Dual Payload Attach FittingDoT Department of TransportationECE Encapsulated Cargo ElementECS Environmental Control SystemEMC Electromagnetic CompatibilityEMI Electromagnetic Interferencefps Feet Per SecondFSPO Flight Safety Project OfficerFSS Flight Safety StationGFO GEOSAT Follow-OnGSE Ground Support EquipmentGTO Geosynchronous Transfer OrbitHz HertzICD Interface Control DocumentKOMPSAT Korean Multi-Purpose SatelliteLbf Pounds, ForceLC Launch ComplexLDS Launch Director StationLEO Low Earth OrbitLEV Launch Equipment VanLSE Launch Support EquipmentLSV Launch Support VanMAB Missile Assembly BuildingMDL Mission Data LoadMIWG Mission Integration Working GroupMSPSP Missile System Pre-Launch Safety

PackageMTI Multi-Thermal Imager

MUX MultiplexerNASA National Aeronautics and Space

Administrationnm Nautical MilesNVAFB North Vandenberg Air Force BaseOASPL Overall Sound Pressure LevelOR Operations RequirementsOSM Operations Support ManagerPDU Pyrotechnic Driver UnitPOST Program to Optimize Simulated Tra-

jectoriesPPF Payload Processing FacilityPRD Program Requirements DocumentPSA Payload Support AreaRCC Range Control CenterRCS Reaction Control SystemRF Radio FrequencyRWG Range Working GroupSLC Space Launch ComplexSPL Sound Pressure LevelSRM Solid Rocket MotorSRS Shock Response SpectrumSSLV Standard Small Launch VehicleSTEX Space Technology ExperimentSVAFB South Vandenberg Air Force BaseTAA Technical Assistance AgreementTCS Test Conductor StationTML Total Mass LossTTL Transitor - Transitor LogicTVC Thrust Vector ControllerTWP Taurus Work PackageU.S. United StatesVAFB Vandenberg Air Force BaseVC-HS Visibly Clean-Highly SensitiveVCS Vehicle Control StationVDC Volts, Direct CurrentVDL Voice Direct LineWBS Work Breakdown StructureWFF Wallops Flight Facility

Section 1Introduction

Release 3.0 September 1999 1-1

Taurus User's Guide Section 1.0—Introduction

This Taurus User’s Guide is intended to familiarizepayload mission planners with the capabilities ofOrbital Sciences Corporation’s (Orbital’s) Tauruslaunch service. This document provides anoverview of the Taurus system design and adescription of the services provided to ourcustomers. Orbital offers a variety of serviceoptions to allow the maximum flexibility insatisfying the objectives of single or multiplepayloads.

The Taurus Program’s philosophy of placingmission success as the highest priority is reflectedin the results of all Taurus missions to date. Thesemissions not only demonstrated the reliability ofthe Taurus launch system design, but also providea solid base from which to conduct our currentTaurus missions.

The Taurus launch vehicle system is composedof a flight vehicle and ground support equipment.Each element of the Taurus system has beendeveloped to simplify the mission design andpayload integration process and to provide safe,reliable space launch services. This User’s Guidedescribes the basic elements of the Taurus systemas well as optional services that are available. Inaddition, this document provides general vehicleperformance, defines payload accommodationsand environments, and outlines the Taurusmission integration process.

The Taurus system can operate from a wide rangeof launch facilities and geographic locations.The system is compatible with, and will typicallyoperate from, existing U.S. Government rangesat Vandenberg Air Force Base (VAFB), CapeCanaveral Air Station (CCAS), and Wallops FlightFacility (WFF). While most commercial missionsdo not require the rapid response capabilities ofthe original Taurus program, the same designfeatures that allow the Taurus system to meetthese requirements are used to streamline vehicleintegration and launch operations for all Taurusmissions. This User's Guide describes Taurus-unique integration and test approaches (includingthe typical operational timeline for payloadintegration with the Taurus vehicle) and theexisting ground support equipment that is used toconduct Taurus operations.

Section 2Description of Taurus

Launch Services


Taurus User's Guide Section 2.0—Description of Taurus Launch Services

2.1 Taurus Launch System Overview

The Taurus launch vehicle, shown in Figure 2-1,was privately developed by Orbital to provide acost effective, reliable and flexible means ofplacing small satellites into orbit. An overview ofthe system and available launch services isprovided within this section, with specificelements covered in greater detail in thesubsequent sections of this User's Guide.

Taurus was originally designed to meet the needsof the Defense Advanced Research ProjectsAgency's (DARPA's) Standard Small LaunchVehicle (SSLV) Program. The requirements ofthat program stressed system reliability,transportability, and operation from austerelaunch sites. Following the successful Taurusmaiden flight, Orbital initiated development workon an upgraded version of the Taurus vehicle.This configuration, which is the Taurus vehiclethat is offered commercially and described withinthis User's Guide, differs from the SSLVconfiguration primarily through the use ofThiokol’s Castor 120® solid rocket motor for thefirst stage in place of the Peacekeeper stage usedon the first flight. Orbital’s efforts have resultedin a simple, robust, self-contained launch systemthat has been completely successful in its threeflights to date and is in production to supportvarious government and commercial customers.The Taurus flight history and current manifest aredepicted in Figure 2-2.

The Taurus system also includes a complete setof transportable Launch Support Equipment (LSE)designed to allow Taurus to be operated as a self-contained satellite delivery system. This LSE ispictorially represented in Figure 2-3 and includesthe launch stand, Launch Equipment Van (LEV),Launch Support Van (LSV) and assortedmechanical and electrical ground supportequipment. The LSV serves as the actual controlcenter for conducting a Taurus launch andincludes consoles for Orbital, range safety, andpayload personnel. While the Taurus system iscapable of self-contained operation, it is typicallylaunched from an established range; thus, thevehicle and LSE are designed for compatibility Figure 2-1. Taurus Vehicle Configuration.

UG-01

Payload Fairing

Avionics Module• Flight Computer• Inertial Navigation System• Telemetry System• Electrical Power• Reaction Control System• Pyro Driver Unit

Stage 3 Assembly• Pegasus Stage 3• Gimballed Nozzle TVC• Flight Termination System

Stage 2 Assembly• Pegasus Stage 2• Gimballed Nozzle TVC• Flight Termination System• Pyro Driver Unit

Stage 1 Assembly • Pegasus Stage 1

• Gimballed Nozzle TVC• Flight Termination System• Pyro Driver Unit

Stage 0 Assembly• Thiokol Castor 120• Gimballed Nozzle TVC• Flight Termination System

Stage 0/1 Interstage



UG-03Figure 2-2. Taurus Flight History and Current Launch Manifest.

DARPA Taurus SSLV (T1)

Commercial Taurus Development

Commercial Geosat Follow-On (GFO) (T2)

Air Force Space Technology

Experiment (STEX) (T3)

Commercial KOMPSAT (T4)

Air Force Multi-Thermal Imager (MTI) (T5)

Commercial OrbView-4 (T6)

Mar 13, 1994

Feb 10, 1998

Oct 3, 1998

Nov, 1999

1st Qtr, 2000

3rd Qtr, 2000

Launch Date 1992 1993 1994 1995 1996 1997 1998 1999 2000 2001 2002 2003 2004Task Name

Successful

Successful

Successful

Figure 2-3. Transportable Launch Support Equipment.UG-05

• Orbital Net• Countdown Net• Safety Net• Countdown Clock• Status/Alerts• Pad Video• VDLs• Prelaunch Telemetry

Launch Support Van

Launch Equipment Van• Vehicle Power • Signal Distribution• Payload Support Equipment Taurus

• Countdown Net• Mission Directors Net• Countdown Clock• Status/Alerts

• Pad Video• VDLs• Orbital Net• Prelaunch Telemetry

Range Control Center



with existing U.S. Government ranges atVandenberg Air Force Base (VAFB), CapeCanaveral Air Station (CCAS) and Wallops FlightFacility (WFF). A communications networkconnects the LSV with the Range Control Center(RCC). Further description of the Launch SupportEquipment is provided in Section 2.4.

2.2 Taurus Launch Service

As summarized in Figure 2-4, Orbital provides allof the necessary hardware, software and servicesto integrate, test and launch a satellite into itsprescribed orbit. In addition, Orbital will completeall the required agreements, licenses anddocumentation to successfully conduct Taurusoperations. All Taurus production and integrationprocesses and procedures have beendemonstrated and are in use for current Taurusmissions. The Taurus mission integration processcompletely identifies, documents, and verifies

all spacecraft and mission requirements. Thisprovides a solid basis for initiating andstreamlining the integration process for futureTaurus customers.

2.3 Description of the Taurus Vehicle

The Taurus vehicle, shown in expanded view inFigure 2-5, is a four stage, inertially guided, allsolid propellant ground launched vehicle.Conservative design margins, state-of-the-artstructural systems, a distributed-processoravionics architecture, and simplified integrationand test capability, yield a robust, highly reliablelaunch vehicle design. In addition, Tauruspayload accommodations and interfaces havebeen designed to satisfy a wide range of potentialpayload requirements.

The Taurus vehicle was designed to enablepayloads to choose the optimum configuration

UG-23

Figure 2-4. Overview of Taurus Launch Service.

Launch System Management And Production• Furnish The Personnel, Facilities, And Materials To Manufac-

ture, Test, Integrate, Support And Launch The Taurus Vehicle• Furnish The Personnel, Facilities, And Materials To Design,

Manufacture, Test, And Integrate Mission-Unique HardwareAnd Software- Electrical, Mechanical And GSE Interfaces

• Perform Standard Mission AnalysesMission Integration• Coordinate The Mission Integration Working Group (MIWG)

Process• Create And Maintain The Taurus-to-Payload Interface Con-

trol Document• Define Mechanical And Electrical Interfaces, Payload En-

vironments, Vehicle Performance Requirements• Conduct The Taurus-to-Payload Interface Verification Process • Support Interface Testing Prior To Payload Transport To The

FieldRange Interface• Coordinate The Mission With Applicable Range And Environ-

mental Agencies• Coordinate, Prepare And Maintain Range Documentation

(UDS) Submittals With Customer-Provided Payload Input• Conduct Flight Planning And Trajectory AnalysesSafety• Implement The Safety Program And Obtain All Necessary

Flight Certification And Range Approvals• Complete The Missile System Pre-Launch Safety Package

(MSPSP), Providing Detailed Technical Data Of All LaunchVehicle And Payload Hazardous Items

• Submit The Flight Data Safety Package For Flight PlanApproval

• Complete The Launch Complex Safety Plan

Launch Base Services• Conduct/Provide All Activities, Supplies and Services At The

Integration Site Necessary To Build And Test Taurus• Create And Maintain Taurus-Unique And Mission-Integrated

Field Site Procedures• Provide Facility For Payload Processing And Encapsulation• Provide Environmental Control Of Payload From Encapsula-

tion Through Launch• Provide Space For Payload Personnel And Equipment In

Orbital Support Vans• Provide Electrical Interface From The Payload Interface

Connectors To The LSE• Transport Components To Launch SiteLaunch Base Integrated Operations• Prepare Payload Adapter, Fairing And Mission Support

Hardware• Perform Encapsulation And Transportation Of Encapsulated

Cargo Element (ECE) To The Launch Site• Conduct Mate Of ECE To Launch Vehicle And Interface

Verification TestingLaunch And Monitor Of Taurus System• Develop And Maintain Mission Constraints Document And

Launch Checklist, Which Defines The Pre-Launch AndLaunch Activities And Procedures

• Conduct Countdown Rehearsals With Full Range And Cus-tomer Participation To Assure Readiness For Launch

• Provide Orbital Parameters (From Taurus Telemetry)• Confirm Payload Separation, Time Of Separation And

Separation State Vector (If Within Sight Of Tracking Station)• Process Flight DataCommercial Services• Obtain And Maintain DoT Launch License• Obtain TAA for Non-U.S. Customers



that best meets mission requrements. In order tosimplify the options offered, Orbital created aconfiguration numbering convention detailed inFigure 2-6. This numbering system is used todescribe the performance of Taurus vehicles inthe following sections.

Stage 0 — The Taurus first stage (known as Stage0) is the Thiokol Castor 120 Solid Rocket Motor(SRM). The flight-proven Castor 120 SRM, is acommercial replacement of the reliablePeacekeeper first stage flown on the Taurusmaiden flight. The Castor 120 physical andperformance characteristics are closely relatedto the Peacekeeper motor. Propellant binder,aluminum and solid content are the same, lengthand weight are increased modestly, and ballisticsare reshaped slightly to create a more benignthrust profile and flight environment whileachieving higher performance. Castor 120 motordimensions, mass and performancecharacteristics are provided in Figure 2-7.

Stage 0/1 Interstage — An aluminum skin andstringer construction interstage extends from the

forward skirt of the Castor 120 Stage 0 motor tothe aft end of the Stage 1 motor, tapering from a92 inch (2.3 m) to a 50 inch (1.3 m) diameter.This interstage is fabricated in two sections: thelower interstage, which remains with Stage 0,and the upper interstage, which flies with Stage1. A field joint between the upper and lowersections of the interstage allows the Taurus upperstage stack to be mated to the Castor 120 Stage 0,while a linear shaped charge separation systemsevers the interstage sections at Stage 1 ignition.The lower interstage is designed to accommodatea hot-fire separation of Stage 1 through theincorporation of vent ports in the cylindricalportion of the structure. Aerodynamic panelscovering the ports are jettisoned just prior toStage 1 ignition. This hot-fire separation hasbeen successfully demonstrated on all Taurusmissions.

Stages 1, 2, and 3 — The Taurus upper stages(known as Stages 1, 2 and 3) are the AlliantTechSystems Orion 50S-G, 50 and 38 SRMs,respectively. These motors were originallydeveloped for Orbital's Pegasus program and

Composites - Orbital/VermontComposites

Fairing 63"92"

- Orbital/Vermont Composites- Orbital/R-Cubed Composites

Motors - Alliant Techsystems

TVC - Allied Signal

TVC - Parker

UG-06

Flight Computer

IMUPyro Driver UnitBatteriesMultiplexerPower Transfer SwitchFTS Receiver TransponderWire Harness

SBS Embedded ComputersLittonOrbitalOrbital, BSTOrbitalOrbital

Cincinnati ElectronicsHerley MicrowaveOrbital

Avionics

TVC - Allied Signal

Motor - Thiokol

Reaction Control System - Moog, Orbital

Stage 2

Stage 3

Interstage - Orbital/Remmele

Figure 2-5. Taurus Expanded View.

Stage 0

Stage 1



have been adapted for use on the ground-launched Taurus vehicle. Common designfeatures, materials and production techniquesare applied to all three motors to maximizereliability and production efficiency. Thesemotors are fully flight qualified based on heritage,design conservatism, ground static fires andmultiple successful flights. Details of Stages 1, 2and 3 are also provided in Figure 2-7.

As an optional service, the Orion 38 Stage 3 canbe replaced by a spinning upper stage usingThiokol’s STAR 37FM motor. This optional serviceis detailed in Section 8.6.

Avionics Section and Boattail — The graphite/epoxy boattail structure provides the transitionfrom the 50 inch (1.3 m) diameter of the Stage 2motor to the 63 inch (1.6 m) diameter of theavionics skirt. The boattail is a sandwich structureconstructed using a foam core and graphite epoxyfacesheets. The avionics skirt, also a graphite/epoxy structure, supports the avionics shelf andcarries the primary structural loads from thefairing and payload cone. The aluminum avionicsshelf supports the third stage avionics. The shelfis annular in shape to maximize the use ofavailable volume and enable the Stage 3 motor tobe suspended from the payload cone through the

shelf. This removes the Stage 3 motor from theprimary structural load path and provides theflexibility for alternate upper stage configurations.

Avionics — The Taurus avionics system is an all-digital distributed processor design. Missionreliability is achieved by the use of simple designs,high reliability components, high design marginsand extensive testing at the box, subsystem, andsystem level. The heart of the avionics system isa multiprocessor, 32-bit flight computermanufactured by SBS Embedded Computers. Theflight computer communicates with all vehiclesubsystems and the LSE using standard RS-422digital serial data links. All avionics on thevehicle feature integral microprocessors toperform local processing and to handlecommunications with the flight computer. ThisRS-422 architecture is central to Taurus' rapidintegration and test, as it allows unit-, stage- andsystem-level testing to be accomplished usingonly personal computers outfitted withcommercially available interface cards.

Payload Fairings — Orbital offers one of twoflight-proven Taurus payload fairings toencapsulate the payload and provide protectionand contamination control during groundhandling, integration operations and flight. Bothfairings are bisector shells constructed of graphite/epoxy facesheets with an aluminum honeycombcore. The Taurus 63" diameter fairing,manufactured by Vermont Composites, wassuccessfully demonstrated on Taurus' first andthird missions while the larger 92" diameterfairing, manufactured by R-Cubed Composites,was successfully demonstrated on the secondTaurus mission. The fairings are further detailedin Section 5.2.

1 -

2 -1 -2 -

1 -3 -

0 -SSLV Taurus (Peace-keeper Stage 0)Commercial Taurus(Castor 120 Stage 0)

63"92"

Orion-38STAR-37

Vehicle Configuration FairingSize

Stage 3Motor

UG-02

Figure 2-6. Taurus Configuration Numbering Convention.

2 2 1 0

<None>

Stage 4Motor

Propulsion DataEngine/MotorManufacturerPropellantLength (ft/m)Diameter (ft/m)Propellant Weight (lb/kg)Average Vacuum Thrust (lbf/kN)Eff. Specific Impulse (sec)Burn Time (sec)

Castor 120ThiokolSolid HTPB41.9/12.87.8/2.4107,408/48,720363,087/1,615277.982.5

Stage 0Orion 50S-GAlliantSolid HTPB28.3/8.64.2/1.326,780/12,147106,000/471285.072.4

Stage 1Orion 50AlliantSolid HTPB10.1/3.14.2 /1.36,667/3,02425,910/115290.275.1

Stage 2Orion 38AlliantSolid HTPB4.4/1.3 3.2 /1.01,697/7707,155/31.8286.768.5

Stage 3

Figure 2-7. Taurus Propulsion Data Table. UG-04



The Taurus fairings have been designed to satisfya range of payload interface requirements. Bothdesigns provide for the off-line encapsulation ofthe payload within the fairing and thetransportation of this encapsulated cargo elementto the launch site for mating to the Taurus vehiclelate in the launch operations flow. The fairingsalso provide for low payload contaminationthrough prudent design and selection of lowcontaminating materials and processes. Eachfairing incorporates acoustic blankets and internalair conditioning ducts to provide more benignpayload environments.

With the addition of a structural adapter, eitherfairing can accommodate multiple payloads. Thisfeature, described in more detail in Section 8.2 ofthis User's Guide, permits two or more smallerpayloads to share the cost of a Taurus launch,resulting in a lower launch cost for each ascompared to other launch options.

2.4 Primary Launch Support Equipment

The Taurus electrical LSE consists of two mainoperational units: the LSV and the LEV. Thesevans house all the electrical equipment requiredto safely and successfully monitor, control, andlaunch the Taurus vehicle. The LSV serves as thecontrol center during the launch countdown andcan accommodate ten personnel: three customer,two range safety, and five Orbital representatives.As shown in Figure 2-8, The LSV is locatedapproximately 5 miles (8.0 km) away from thelaunch site. The unmanned LEV, locatedapproximately 200 ft (61 m) from the launch pad,acts as a junction equipment room patching thefiber optic lines from the LSV into a copperconnection with the vehicle. Both vans aresimilar in construction, consisting of corrugatedsteel shelters equipped with air conditioning,dual personnel exits, equipment loading doors,and interior lighting. Both shelters are prewired

LEV

UG100A

Figure 2-8. Taurus LSE Configuration.

CopperWires

(Vehicleto LEV)

Fiber Optic(LEV to LSV)

≤5 Miles(≤8 km)

200 Feet(61 m)

LSV

Launch Vehicle

ECS



to support up to 200 A of single phase 208Vpower.

Mission-unique, customer-supplied payloadconsoles and equipment can be supported in theLSV and LEV. Interface to the payload throughthe Taurus T-0 umbilicals and land lines providesthe capability for direct monitoring of payloadfunctions.

2.4.1 Launch Support Van

The LSV, shown in Figure 2-9, consists of a 10 ftx 40 ft (3.1 m x 12.2 m) shelter housing all thehardware required to monitor and control thevehicle and payload. This hardware includesfour consoles (one control and three monitorconsoles), a payload support area, and a singlethree-bay equipment rack for interfacing to theLEV and the range. The four consoles havededicated functions and are differentiated asfollows: Flight Safety Station (FSS), VehicleControl Station (VCS), Test Conductor Station(TCS), and the Launch Director Station (LDS).

The LSV also contains a telephone service, acentrally located countdown clock andcommunications and telemetry interfaces forconnection with the RCC. All stations haveuninterruptable power supplies.

Flight Safety Station — The FSS is divided intotwo sections, the Flight Safety Project Officer(FSPO) section and the Operations SupportManager (OSM) section. The FSPO sectiondisplays vehicle telemetry for monitoring vehiclesafety, while the OSM section is dedicated tomonitoring pad operations and includes dualClosed Circuit Television (CCTV) monitors. Bothconsoles maintain intercom and telemetrycommunication with the Range Safety Officer inthe RCC.

Vehicle Control Station — The VCS is the primarycontrol station for the vehicle and the LEV. Orbitalengineers at this station monitor vehicle healthand status and perform critical vehicle operationsduring countdown activities. The VCS also houses

UG101

Figure 2-9. LSV Layout.

FlightSafetyStation

VehicleControlStation

TestConductor

Station

LaunchDirectorStation

PayloadSupport

Area

9.6 ft (2.9 m)

Crosshatched Area Designates

Payload Support Area

10 ft

(3.1 m)

40 ft (12.2 m)



UG102

Figure 2-10. LEV Layout.

7.6 ft (2.3 m)

20 ft (6.1 m)

8 ft

(2.4 m)

5.8 ft (1.8 m)

the manual switch panel which gives the operatormanual vehicle safing and hold-fire capability.Two CCTV monitors are located above theconsole for viewing pad activities.

Test Conductor Station — The TCS is a monitor-only station from which the Test Conductorcontrols the final count activities.

Launch Director Station — The LDS is anothermonitor-only station where the customer andOrbital launch directors are located duringlaunch. Each position is equipped with thecapability of monitoring vehicle telemetry andcommunicating with the RCC. The customerposition also has the capability of monitoringpayload specific telemetry.

Payload Support Area (PSA) — The PSA is anarea for payload use capable of supporting twopayload personnel. Further detail is provided inSection 5.4.11.

2.4.2 Launch Equipment Van

The LEV, shown in Figure 2-10, is an 8 ft X 20 ft(2.4 m x 6.1 m) shelter that houses the interface

between the LSV and the vehicle. This unmannedfacility is located approximately 200 ft (61 m)from the launch pad and serves as a junctionequipment room patching fiber optic lines fromthe LSV into a copper connection with the vehicle.The three-bay equipment rack within the LEVhouses equipment required for ground commandand checkout of the vehicle, including powersupplies for battery charging, analog to digitalconverters for monitoring the vehicle analogstatus, and switching logic for commandingvehicle states. Equipment is provided in the LEVto allow complete communications with the LSVand launch site.

Section 3General Performance


Taurus User's Guide Section 3.0—General Performance

3.1 Vehicle Performance

This section describes the orbital performancecapabilities of the Taurus vehicle. Taurus is wellsuited for Low Earth Orbit (LEO) missions to awide variety of altitudes. Taurus can attain arange of posigrade and retrograde inclinationsthrough the choice of launch sites made availableby the transportable nature of the Taurus launchsystem. High energy and GeosynchronousTransfer Orbit (GTO) missions can also beachieved with the standard configuration.Performance to these orbits can be increasedsubstantially using the performanceenhancements discussed in Section 8.6.

3.2 Mission Description

Orbital provides each Taurus launch servicecustomer a mission design optimized to meetcritical requirements while satisfying payload,launch vehicle, and range safety constraints.Orbital uses the Program to Optimize SimulatedTrajectories (POST) to provide projected Taurusperformance values to candidate orbits. POSTcan be set to optimize for mass to a specifiedorbit, for the resultant orbit given a specifiedpayload mass, or for a wide variety of otherconditions. Specific orbit parameters, tolerances,and performance requirements for the missionare defined in the payload Interface ControlDocument (ICD).

There are several different types of trajectoriesthat are used with the Taurus launch vehicle.Since Taurus uses solid rocket motors, theincrements of velocity change are relatively fixed,driving the trajectory design based on the type oforbit that is required.

For LEO missions, Stages 0, 1 and 2 are burned inrapid succession to place the payload and upperstage in a suborbital trajectory with an apogeealtitude equal to the perigee of the final orbit. Thecombined upper stage and payload coast up tothis altitude, where the Stage 3 burn adds thevelocity required to place the payload in itsdesired orbit.

There are certain spacecraft orbits where theposition of apogee and perigee are constrained.

Examples include GTO where perigee must be atthe equator, Molniya orbits where the apogee ispreferred over a certain Earth latitude, and escapetrajectories where the escape velocity vectormust satisfy direction constraints in theheliocentric reference frame. The conventionalapproach to these trajectories is to first insert thepayload and upper stage into a parking orbit.When the vehicle reaches the desired perigeelocation, the upper stage is fired, putting thepayload into its final orbit. Another techniquecan be used to further improve performance tothese orbits. Orbital refers to this technique as“non-apsidal injection” since the upper stagemotor fires at a point that is neither apogee norperigee. The non-apsidal injection techniqueplaces the payload and upper stage into asuborbital "parking orbit" (perigee altitude <0nm), which allows a heavier mass to be lofted toa lower energy state. The upper stage is ignitedafter the apogee of the suborbital trajectory toplace perigee at the desired location. Thistechnique can result in performance gains ofgreater than 25% over a conventional parkingorbit technique to certain orbits, but it imposesmore constraints on available apsidal alignment.Orbital will work with the customer to design thetrajectory that best fits the needs of the payload.

3.3 Taurus Ascent Profile

Orbital shapes the trajectory flown by the Taurusvehicle to realize maximum performance withoutviolating any vehicle design constraints (such asstructural limits) or payload requirements (suchas free molecular heating). This shaping isaccomplished by controlling the product ofdynamic pressure (Q) and angle of attack (α).The Taurus mission profile will vary significantlydepending on the requirements of the payload.The following sections present typical missionprofiles for a direct ascent mission of a standardTaurus 2210 vehicle to a 300 nm (556 km)circular sun-synchronous orbit, and a GTOmission using a Taurus 2110 vehicle and a 150nm x 165 nm (278 km x 306 km) parking orbit.This information is for reference only and willvary on a mission-to-mission basis.



3.3.1 Direct Ascent

A Taurus direct ascent trajectory from a SouthVAFB launch site to a 300 nm (556 km) circularorbit inclined at 97.6° (sun-synchronous) isdepicted in Figure 3-1. Following Stage 0 motorignition, the vehicle performs a short vertical riseto clear the umbilical tower and then performs apitch-over. This sequence, along with the rest ofthe Stage 0 flight profile, is performed by open-loop guidance governed by inertial steering tables.The open-loop steering tables are the result of aQ-α profile design driven by mission requirementsand vehicle structural limits. At 75 seconds thevehicle assumes a zero angle of attack attitude forthe Stage 0/1 separation event, which nominallyoccurs at 81 seconds. The staging event isinitialized as the vehicle acceleration drops to0.2 g’s (corresponding to approximately 10,000lbf [45 kN] of Stage 0 thrust). At this point theStage 1 motor is ignited and, 95 msec later, thetwo stages are pyrotechnically separated. Thissequence of events maintains acceptable controlauthority throughout the staging event.

Stage 1 guidance, like Stage 0, utilizes an openloop command table where the Taurus vehicleflies a pre-programmed pitch and yaw profile.The Stage 1 motor nominally burns for 76 seconds,during which time yaw steering, if required forrange safety or orbital inclination reasons, isutilized. The coast time between Stage 1 burnoutand Stage 2 ignition is typically 12 seconds,although it can be longer for low altitude missions.These missions may require more time to gainaltitude and reduce free molecular heating priorto payload fairing separation. Nominally, thefairing is jettisoned three to five seconds intoStage 2 burn, after the free molecular heating hasfallen below 360 BTU/ft2/hr (1134 W/m2).

During its 78 second burn, Stage 2 uses anexplicit guidance routine that calculates thedesired thrust attitude based on the measuredvehicle energy state compared against the desiredtargets. This guidance scheme improvesperformance and reduces injection errors.Nominally, the Stage 2 burnout condition leavesthe vehicle on a suborbital path peaking slightly

Stage 0 Ignition & LiftoffT (Time from Lift-Off) = 0 sh (Altitude) = 0 nmiV (Velocity) = 0 fpsR (Range) = 0 nmi

Stage 0 Burnout,Stage 0/1 Separation& Stage 1 IgnitionT = 81 sh = 26 nmiV = 6,894 fpsR = 27 nmi

Stage 2 IgnitionT = 170 sh = 86 nmiV = 14,596 fpsR = 169 nmi

Fairing SeparationT = 173 sh = 88 nmiV = 14,732 fpsR = 176 nmi

Stage 2 BurnoutT = 248 sh = 138 nmiV = 21,721 fpsR = 384 nmi

Stage 2/3 Coast

Stage 1 Burnout/SeparationT = 158 sh = 78 nmiV = 14,698 fpsR = 142 nmi

Stage 2 Separation& Stage 3 IgnitionT = 666 sh = 299 nmiV* = 19,964 fpsR = 1,698 nmi

Stage 3 Burnout/Orbit InsertionT = 735 sh = 300 nmiV* = 24,875 fpsR = 1,934 nmi

Payload SeparationT = 795 s

Figure 3-1. Typical Taurus Direct Ascent Mission Profile.UG29

* Inertial Velocity



above the desired mission altitude (or perigeealtitude for an elliptical orbit).

During the coast to the final burn point, theexplicit guidance routine continues to operate,calculating the desired Stage 3 ignition pointbased on measured orbital parameters. TheStage 3 burn, which nominally lasts 69 seconds,provides injection into the final orbit.

3.3.2 Parking Orbit Ascent

The sequence of events for a parking orbit missionis shown in Figure 3-2. Stages 0 and 1 fly amission profile similar to that of a direct ascentmission. Following Stage 1 burnout, a long coastallows the payload and remaining stages to reachthe desired parking orbit injection point. TheStage 2 motor then ignites, placing the payloadand Stage 3 in the parking orbit. The spent Stage2 motor is jettisoned, leaving the Taurus avionicssection and Stage 3 motor attached to the payload.Another long coast occurs as the guidance systemwaits for the appropriate true anomaly value toignite Stage 3 and achieve the correct payload

apsidal targets. The total mission time isconstrained by vehicle power and thermallimitations, so the feasibility of this approachmust be assessed on a mission-specific basis.

3.4 LEO Performance Capability

Taurus performance curves for circular orbits ofvarious altitudes and inclinations are detailed inFigure 3-3 through Figure 3-8 for launches fromCCAS and VAFB. These performance curvesprovide the total mass available to the payload—the mass of the Orbital-provided separation systemhas been accounted for in the Taurus massproperties. Note that the VAFB performancecurves are provided for two launch sites: theexisting Taurus Space Launch Complex 576E(SLC 576E) on North Base and a proposed SouthBase location. South Base launches do nottypically require yaw steering to meet rangesafety over-flight constraints; hence, aperformance benefit is realized. Figure 3-9shows the available launch azimuths from VAFBand CCAS. Figure 3-10 illustrates the stage

Stage 0 Ignition & LiftoffT (Time from Lift-Off) = 0 sh (Altitude) = 0 nmiV (Velocity) = 0 fpsR (Range) = 0 nmi

Stage 0 Burnout,Stage 0/1 Separation& Stage 1 IgnitionT = 81 sh = 22 nmiV = 7,375 fpsR = 34 nmi

Stage 1 Separation/Stage 2 IgnitionT = 363 sh = 149 nmiV* = 16,514 fpsR = 662 nmi

Fairing SeparationT = 366 sh = 150 nmiV* = 16,693 fpsR = 670 nmi

Stage 2 BurnoutT = 441 sh = 152 nmiV*= 25,571 fpsR = 898 nmi

Stage 2/3 CoastParkingOrbit

Stage 1 Burnout/SeparationT = 158 sh = 67 nmiV = 16,088 fpsR = 165 nmi

Stage 2 Separaton& Stage 3 IgnitionT = 1,415 sh = 156 nmiV* = 25,390 fpsR = 4,580 nmi

Stage 3 Burnout/Orbit InsertionT = 1484 sh = 159 nmiV* = 33,323 fpsR = 4,883 nmi

Payload SeparationT = 1,544 s

Figure 3-2. Typical Taurus Parking Orbit Mission Profile.UG30

* Inertial Velocity



Taurus 2110 VehicleTaurus 2210 Vehicle

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Figure 3-3. Taurus Performance to 28.5° Orbits - CCAS Launch.UG23



Figure 3-4. Taurus Performance to 38° Orbits - CCAS Launch.UG24


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Figure 3-5. Taurus Performance to 60° Orbits - South VAFB Launch.UG25

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Figure 3-6. Taurus Performance to 70° Orbits - South VAFB Launch.UG26


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Figure 3-7. Taurus Performance to 90° Orbits - VAFB Launch.UG27


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Taurus 2110 Vehicle - SVAFBTaurus 2110 Vehicle - NVAFBTaurus 2210 Vehicle - SVAFBTaurus 2210 Vehicle - NVAFB

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Figure 3-8. Taurus Performance to Sun Synchronous Orbits - VAFB Launch.UG28

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vacuum impact points for the two trajectoriesprofiled in Section 3.3.

For missions that require high inclination orbits(greater than 55°), launches are conducted fromVAFB. Orbital established SLC 576E on NorthVAFB for the first several Taurus flights and willhave this facility available for future West Coast

Taurus launches. The minimum launch azimuthacceptable from SLC 576E is 205°, requiring yawsteering (dogleg trajectories) for inclinations belowapproximately 105°. For intermediate inclinations(down to 55°), Orbital can utilize a launch site onSouth VAFB to avoid the performance penaltyincurred by yaw steering. The effect of yaw

Figure 3-10. Typical Stage Vacuum Impact Points for Taurus Launches from VAFB and CCAS.UG-31

Figure 3-9. Launch Range and Site Selection Are Tailored to Mission Performance and Safety Require-ments.

UG-16

VAFB-Western Range• High-Inclination Orbits• Polar and Sun-Synchronous Orbits

CCAS-Eastern Range • Low-Inclination Orbits • Planetary (Escape) Trajectories

Launch Range Capabilities

°

°

°

110

576E/SVAFBOrbit Inclinations

55° – 120°(Nominal)

LC-46Orbit Inclinations

28.5° – 40°(Nominal)

158

220

FlightAzimuths

60

°

FlightAzimuths



steering for range safety reasons is includedwhere appropriate.

Orbital will use CCAS as the launch site formissions to orbital inclinations between 28.5°and 40°.

The NASA WFF launch site can be used formissions requiring intermediate orbitalinclinations (38° - 55°). Southeast launches fromWFF offer fewer overflight concerns than CCASand inclinations up to 55° are feasible withdoglegs and altitude constraints due to stageimpact considerations.

3.5 Injection Accuracy

Orbital injection errors for Taurus are dominatedby variations in the Stage 3 total impulse,navigation errors, and other stage performanceparameters. Taurus 3-σ injection accuracies aresummarized in Figure 3-11.

3.6 Payload Deployment Attitude Options

Following orbit insertion, the Taurus Stage 3avionics subsystem can execute a series of pre-programmed RCS maneuvers to provide thedesired initial payload attitude prior to separation.This capability may also be used to incrementallyreorient Stage 3 for the deployment of multiplespacecraft with independent attituderequirements. Either an inertially-fixed or spin-stabilized attitude may be specified by thecustomer.

The maximum spin rate for a specific missiondepends upon the spin axis moment of inertia ofthe payload and the amount of RCS propellantneeded for other attitude maneuvers, but cannotexceed 15 rpm. Figure 3-12 provides the typical

payload pointing and spin rate accuracies,although these will vary with payload massproperties.

3.7 Separation Tip-Off Rates and Velocities

Orbital performs preliminary and final mission-specific tip-off analyses for each payload, usingthe commercial Dynamic Analysis and DesignSystem (DADS) software. These simulationsverify that spacecraft tip-off and separationvelocity requirements are met even under 3-σconditions. Payload tip-off refers to the angularvelocity imparted to the payload upon separationdue to an uneven distribution of torques andforces. Tip-off rates are highly dependent onpayload mass properties and geometry, but aretypically on the order of 1°/sec.

Separation velocities are driven by the need toprevent recontact between the payload and theTaurus upper stage after separation. The actualvalue will vary with mass properties, but willtypically be approximately 2 ft/sec (0.6 m/sec).

Error Type Angle Rate

3-Axis

Spinning

YawPitchRoll

Spin AxisSpin Rate

±0.7°±0.7°±0.7°±1.0°N/A

±0.4°/sec±0.4°/sec±0.4°/sec≤15 rpm±1.5°/sec

Figure 3-12. Typical Pre-Separation Payload Point-ing and Spin Rate Accuracies.

UG

-19

Injection ApseNon-Injection Apse

Mean AltitudeInclination

±5 nm (±10 km)±27 nm (±50 km)±16 nm (±30 km)

±0.15°Figure 3-11. Injection Accuracies to Low Earth Or-bits (3-σ).

UG

-18

Error Type Tolerance

Section 4Payload Environments


Taurus User's Guide Section 4.0—Payload Environments

4.1 Payload Environments

This section provides details of the predictedenvironmental conditions that the payload willexperience during Taurus ground operations,powered flight, and orbital operations.

Taurus ground operations include payloadencapsulation within the fairing, subsequenttransportation to the launch site and vehicleintegration activities. Powered flight begins atStage 0 ignition and ends at Stage 3 burnout.Taurus on-orbit operations begin after Stage 3burnout and end following payload separation.In order to better define simultaneous conditions,the powered flight portion of the mission isfurther subdivided into smaller time segmentsbounded by critical flight events such as motorignition, stage separation, and transoniccrossover.

The environmental design and test criteriapresented in the following section have beenderived using measured data obtained fromprevious Taurus missions, motor static fire tests,and other system development tests. These criteriaare applicable to Taurus configurations usingeither the 63" or 92" diameter fairings. Thepredicted levels presented are intended to berepresentative of mission specific levels. Missionspecific analyses are performed as a standardservice and are documented in the mission ICD.

Dynamic loading events that occur throughoutthe flight include steady state acceleration,transient low frequency acceleration, acousticloading, sinusoidal and random vibration, andpyrotechnic shock events. Figure 4-1 identifiesthe time phasing of these dynamic loading eventsand their significance. Pyroshock events are notindicated as they do not occur simultaneous withany other significant dynamic loading events.

4.2 Steady State and Transient Accelerations

Design limit load factors due to the combinedeffects of steady state, low frequency transientand quasi-sinusoidal accelerations are defined inFigure 4-2. These values include uncertaintymargins. Note that loads produced by pressureoscillations generated during solid rocket motorcombustion, known as resonant burn, produce aresponse that is both transient and quasi-sinusoidal in nature.

Lateral lift-off transients are generated by windloading at launch stand separation augmentedby vortex shedding, and asymmetric ignitionover pressure. Axial lift-off transients result frompressure oscillations at motor ignition. Duringthe subsonic portion of flight, the maximum axialaccelerations generally occur as a result ofresonant burn excitation. Steady state lateralaccelerations are relatively benign at this point inthe flight. Near transonic crossover, gust andbuffet induced lateral loads are assumed to actsimultaneously with resonant burn induced axialloads. Later in the ascent, as the vehicleapproaches maximum dynamic pressure (MaxQ), buffeting becomes less severe although gustloading continues to contribute to lateral loads.Steady state axial accelerations continue to riseduring Stage 0 ascent finally peaking at the endof burn. Toward the end of Stage 0 burn, transientaccelerations are negligible. The maximumsteady state accelerations occur at the end ofStage 1 burn. At this point in the ascent, the onlyother significant dynamic event occurring isresonant burn excitation. During upper stagemotor ignition, burnout, and staging events,transient loads are relatively benign.

As dynamic response is largely governed bypayload characteristics, mission specific coupled

UG-032Figure 4-1. Phasing of Dynamic Loading Events.

Typical Flight Time (Sec)Steady State LoadsTransient LoadsSinusoidal VibrationAcoustic LoadsRandom Vibration

0-3YesYesYesYesYes

Liftoff4-14YesYesYesNoNo

Subsonic15-25YesYesYes

6 dB Down6 dB Down

Transonic25-45YesYesNoYesYes

Max. Q45-75YesNoNo

6 dB Down6 dB Down

AeroacousticTail-Off

80-155YesYesYesNo

Negligible

S1Burn

157-232YesNoNoNo

Negligible

S2Burn

650-725YesNoNoNo

Negligible

S3BurnTask Name



loads analyses will be performed in order toprovide more precise load predictions. Resultswill be documented in the mission specific ICD.

4.3 Sine Vibration

The Design Limit Load Factors presented in Figure4-2 are suitable for static load purposes; however,they do not address possible fatigue associatedwith lift-off and resonant-burn induced quasi-sinusoidal response in the payload. The customermust determine whether sine vibration testing isrequired or not. Transient loads analysis, andacoustic, random vibration or other tests mayprovide suitable validation of the payload designintegrity and/or workmanship such that sinetesting is not required. In some cases, sufficientmargin against the design limit loads may beavailable to argue that cycling at peak levels doesnot present a credible failure mode.

There are two means to address this quasi-sinusoidal excitation. The classical approach isto define this excitation in terms of a sine sweepenvironment. Figure 4-3 defines the maximumflight level payload interface sinusoidal vibrationlevels associated with the lift-off (10-25 Hz),Stage 0 resonant burn (45-65 Hz) and Stage 1resonant burn (65-75 Hz) forcing functions. Alsoindicated on Figure 4-3 is the lower bound on therange of responses predicted for Taurus payloadsto date. The advantage of the sine sweep approachis that it ensures that all the spacecraft resonanceswithin the band of excitation are stimulated.

An alternative means of addressing the quasi-sinusoidal content is by means of a transient

shaker test. In this approach, the predictedresponse is defined in terms of a low frequencyshock response spectrum (SRS). Figure 4-4 definesthe maximum flight level payload interface quasi-sinusoidal vibration levels in terms of an SRS.The event duration is relatively short in this caseconsistent with the brief period over which theexcitation frequencies and the payload responsefrequencies coincide as determined from coupledloads analysis. Also indicated in Figure 4-4 is thelower bound on the range of responses predictedfor Taurus payloads to date. The advantage ofthis approach is that it provides exposure to amore realistic number of cycles at peak levels.

UG-096a

Figure 4-3. Range of Sine Vibration Levels.

0.8

0.7

0.6

0.5

0.4

0.3

0.2

0.1

0

Leve

l (P

eak

g)

Frequency (Hz)

Payload Interface Lateral MPE Sine Vibration Levels

Upper BoundLower Bound

0 2010 30 50 7040 60 80

Frequency(Hz)

10-2545-6565-75

1.24.51.2

Upper Bound(Peak g)

0.50.70.4

4 Oct/Min24 Hz/Min12 Hz/Min

Lower Bound(Peak g) Sweep Rate

Axial Sine Sweep Specification

Frequency(Hz)

10-2545-6565-75

0.20.50.3

Upper Bound(Peak g)

0.050.1

0.06

4 Oct/Min24 Hz/Min12 Hz/Min

Lower Bound(Peak g) Sweep Rate

Lateral Sine Sweep Specification

54.5

43.5

32.5

21.5

10.5

00 20 40 60 80

Leve

l (P

eak

g)

Frequency (Hz)

Payload Interface Axial MPE Sine Vibration Levels


Event

UG

34

Lift-Off

Subsonic Resonant BurnTransonic

Supersonic Gust

S1 ResonantBurnMotor Burn Out

Axial (g's) Steady State ±Oscillating

(Resultant)-2.5±3.5

(-6.0/+1.0)-3.0±4.0

(-7.0/+1.0)-3.5±4.5

(-8.0/+1.0)-4.0±0.5

(-4.5/-3.5)-4.0±2.0

(-6.0/-2.0)-7.2

Lateral (g's)±2.0

±2.0

±2.5

±2.5

±1.5

±0.5

Figure 4-2. Payload Design C.G. Limit Load Factors.



4.4 Natural Frequency Requirements

In order to minimize loads and deflections aswell as the potential for coupling with the launchvehicle guidance system, the first bendingfrequency of the payload assuming a fixed baseshould be maintained above 25 Hz. If thepayload’s first bending frequency falls below 25Hz, payload lateral loads may exceed thosedefined in Figure 4-2.

Furthermore, in order to minimize loads withinthe payload due to lift-off and resonant burnexcitation, it is recommended that critical axialfrequencies be specified such that the coupledvehicle/payload system frequencies lie between35 and 45 Hz or exceed 75 Hz. If this guidelinecannot be accommodated, payload accelerationloads may exceed those defined in Figure 4-2,Figure 4-3 and Figure 4-4.

Orbital understands that these requirements maybe difficult to meet and will work with thecustomer to recommend practical solutionstoward addressing these issues. Should the needarise, Orbital can provide a spacecraft isolationsystem as a non-standard service (as described inSection 8.5) in order to mitigate the effects ofresonant burn excitation. Such a system wassuccessfully flown twice on Taurus in 1998.

4.5 Random Vibration

The worst-case payload random vibrationenvironment is created by acoustic noisegenerated during lift-off and during ascent nearthe point of maximum dynamic pressure. Thetransmission of this energy and its conversioninto a vibration response at the payload interfaceis highly sensitive to spacecraft mass, geometryand structural design, and thus must bedetermined on a mission-by-mission basis. Figure4-5 defines the maximum flight level payloadinterface random vibration levels as well as thelower bound on responses predicted for Tauruspayloads to date. Also indicated is an envelopeof the flight measured data.

The large spread in predicted response providesan indication of how sensitive the response levelsare to payload configuration. The spectra definedin Figure 4-5 are provided for reference only.They are to be used to evaluate componentsmounted in the vicinity of the payload interfaceand low to mid frequency structure-bornevibration into the entire payload. In order topredict the payload interface response levels,mission-specific analyses will be conducted byOrbital as a standard service using a customerprovided statistical energy model of the payload.

10

1

0.1

Leve

l (P

eak

g)


Frequency (Hz)10 100

Payload Interface Axial MPE Transient ShockResponse Spectrum (SRS/Q)

1

0.1

0.01

Leve

l (P

eak

g)


Frequency (Hz)10 100

Payload Interface Lateral MPE Transient ShockResponse Spectrum (SRS/Q)

Frequency(Hz)

102030456580

100

0.20.20.20.50.50.30.3

Upper Bound(Peak g)

0.050.050.05

0.10.1

0.060.06

Lower Bound(Peak g)

Frequency(Hz)

102030456580

100

0.151.21.24.54.51.21.2

Upper Bound(Peak g)

0.10.50.50.70.70.40.4

Lower Bound(Peak g)

UG-095

Figure 4-4. Range of Shock Response Spectrums.NOTE: Assumes 20 Cycles at Peak Levels



4.6 Acoustic Environment

Fairing interior noise levels for the 63" and 92"diameter fairings are defined in Figure 4-6 andFigure 4-7, respectively. Levels are shown witha maximum fill factor included and with no fillfactor included. Taurus flight microphone datashows that peak acoustic environments occur atlift-off and near the point of maximum dynamicpressure. As noted in the figures, acoustic levelsare dependent on spacecraft geometry (fill factor)and, to a lesser extent, upon acoustic absorptioncharacteristics. As a standard service, mission-specific analyses will be conducted by Orbitalusing a customer provided statistical energy modelof the payload.

4.7 Pyro Shock

The maximum predicted shock levels at thepayload interface are presented in Figure 4-8.These levels are applicable to a payload usingTaurus' standard payload separation system.

4.8 Thermal and Humidity

Ground Operations — Upon encapsulationwithin the fairing and for the remainder of groundoperations, the payload environment will bemaintained by the Taurus Environmental ControlSystem (ECS). Fairing inlet conditions are selectedby the customer, and are bounded as follows:

• Dry Bulb Temperature: 55-85°F (13-29°C)controllable to ±3°F (±2°C) of setpoint;

• Dew Point Temperature: 38-62°F (3-17°C);• Relative Humidity: determined by drybulb

and dewpoint temperature selections and

UG37a

Figure 4-5. Payload Interface Random Vibration Levels.

Frequency (Hz)

PS

D L

evel

(g2

/Hz)

1.00E+00

1.00E-01

1.00E-02

1.00E-03

1.00E-0410 100 1,000 10,000

Frequency

20100200800

1,6002,000Grms

3.00E-044.00E-034.00E-031.20E-011.20E-016.00E-02

12.6

PSD Level(g2/Hz) Frequency

2050

400800

1,2002,000Grms

3.00E-041.20E-031.20E-033.20E-033.20E-031.00E-03

2.0

PSD Level(g2/Hz)

Upper Bound Levels Lower Bound Levels

Frequency

2080

1401,4002,000Grms

1.40E-043.00E-042.40E-031.00E-012.00E-02

9.0

PSD Level(g2/Hz)

Flight Data Envelope

Upper Bound Level (12.6 grms)Flight Data Envelope +3dB (9.0 grms)Lower Bound Level (2.0 grms)

120

Figure 4-6. 63" Diameter Fairing Maximum Flight Level Payload Acoustic Environment.

130

125

115

110

105

100

95

90

85

2031

.5 50 80 125

200

315

500

800

1250

2000

3150

5000

8000

Frequency (Hz)

1/3

Oct

ave

Ban

d S

PL

(dB

) re

: 20

µP

a

MPE Noise Level Assuming Max Fill Factor (136.1dB OASPL)MPE Noise Level Assuming No Fill Factor (133.3 dB OASPL)

Frequency

(Hz)

UG

38a

20

25

31.5

40

50

63

80

100

125

160

200

250

315

400

Max Fill

Factor

1/3 Oct

Band SPL

(dB)Frequency

(Hz)

Max Fill

Factor

1/3 Oct

Band SPL

(dB)

116.1

122.1

122.1

119.6

120.9

125.0

123.4

123.7

123.4

121.4

123.3

122.7

122.7

123.3

Min Fill

Factor

1/3 Oct

Band

SPL (dB)

Min Fill

Factor

1/3 Oct

Band

SPL (dB)

111.5

117.5

117.6

115.1

116.5

120.6

119.1

119.5

119.3

117.5

119.6

119.1

119.4

120.2

500

630

800

1,000

1,250

1,600

2,000

2,500

3,150

4,000

5,000

6,300

8,000

10,000

124.3

122.6

117.8

115.3

119.9

117.8

118.4

117.7

119.7

126.8

122.6

110.5

98.0

90.5

136.1

121.5

120.0

115.5

113.3

118.2

116.3

117.1

116.6

118.7

126.0

122.0

110.1

97.8

90.4

133.3Duration 1 Minute 1 Minute Overall SPL



generally controlled to within ±3%. Relativehumidity is bound by the psychrometric chartand will be controlled such that the dew pointwithin the fairing is never reached.

Powered Flight — The maximum fairing insidewall temperature will be maintained at less than250°F (121°C), with an emissivity of 0.92. Thistemperature limit envelopes the maximumtemperature of any component inside the payloadfairing with a view factor to the payload with theexception of the Stage 3 motor.

The maximum upper stage motor surfacetemperature exposed to the payload will notexceed 350°F (177°C), assuming no shieldingbetween the aft end of the payload and theforward dome of the motor assembly. Whetherthis temperature is attained prior to payloadseparation is dependent upon mission timeline.Fairing deployment will be initiated such that the

maximum free molecular heating rate is less than360 BTU/ft2/hr (1134 W/m2).

4.9 Contamination Control

The Taurus vehicle, all payload integrationprocedures, and Orbital’s contamination controlprogram have been designed to minimize thepayload’s exposure to contamination from thetime the payload arrives at the payload processingfacility through orbit insertion and separation.

The payload processing area will be in a FED-STD-209 Class M6.5 (100,000) clean roomenvironment. The clean room and anteroom(s)utilize HEPA filter units to filter the air. Orbitalwill provide the necessary clean room garments.

The Taurus assemblies that affect cleanlinesswithin the encapsulated payload volume includethe fairing, the acoustic blankets and the payloadcone assembly. The inner surface of theseassemblies and the blanket exterior surface areall cleaned and verified to Visibly Clean HighlySensitive (VC-HS) per JSC-SN-C-0005.

Orbital provides an ECS from payloadencapsulation through vehicle lift-off. The ECScontinuously purges the fairing volume withclean filtered air. Orbital’s ECS incorporates aHEPA filter unit to provide FED-STD-209 ClassM5.5 (10,000) air. Orbital monitors the supplyair for particulates via a probe installed upstreamof the fairing inlet duct.

UG39

Figure 4-8. Maximum Flight Level Pyroshock at thePayload Interface.

10,000

1,000

100

100 1,000 10,00010

Frequency (Hz)

SR

S L

evel

(g'

s)

Interface Shock Levels

Frequency100

1,30010,000

SRS Level(g's)

503,5003,500

120

UG

39a

Figure 4-7. 92" Diameter Fairing Maximum Flight LevelPayload Acoustic Environment.

130

125

115

110

105

100

95

90

85

2031

.5 50 80 125

200

315

500

800

1250

2000

3150

5000

8000

Frequency (Hz)

1/3

Oct

ave

Ban

d S

PL

(dB

) re

: 20

µP

a

MPE Noise Level Assuming Max Fill Factor (138.0 dB OASPL)MPE Noise Level Assuming No Fill Factor (134.0 dB OASPL)

Frequency

(Hz)

20

25

31.5

40

50

63

80

100

125

160

200

250

315

400

Max Fill

Factor

1/3 Oct

Band SPL

(dB)Frequency

(Hz)

Max Fill

Factor

1/3 Oct

Band SPL

(dB)

120.5

126.0

125.9

123.2

124.5

128.5

126.9

127.0

126.2

123.8

125.7

124.7

125.1

126.3

Min Fill

Factor

1/3 Oct

Band

SPL (dB)

Min Fill

Factor

1/3 Oct

Band

SPL (dB)

115.1

120.7

120.6

118.0

119.4

123.5

122.0

122.3

121.7

119.5

121.6

120.9

121.5

123.0

500

630

800

1,000

1,250

1,600

2,000

2,500

3,150

4,000

5,000

6,300

8,000

10,000

125.2

122.7

116.6

114.7

118.8

116.3

116.1

115.2

116.4

122.9

118.9

106.5

98.3

93.2

138.0

122.2

120.1

114.3

112.7

117.0

114.8

114.8

114.1

115.4

122.2

118.3

106.0

97.9

92.9

134.0Duration 1 Minute 1 Minute Overall SPL



The fairing and payload cone assemblies aregraphite reinforced epoxy composite structureswith a total mass loss (TML) value of approximately0.54% by weight and collected volatilecondensable material (CVCM) value ofapproximately 0.02%. Other materials usedwithin the payload encapsulated environmenthave been reviewed for their outgassingcharacteristics, and either have a outgassingcharacteristic of less than 1.0% TML and lessthan 0.1% CVCM or are identified within Taurus'Outgassing Data Report.

4.10 RF Environment

Ascent Environment — As shown in Figure 4-9,Taurus possesses three RF sources: two S-Bandtransmitters at 2288.5 MHz and 2269.5 MHz,respectively, and a C-Band Transponder whichtransmits at 5765 MHz. The payload fairingattenuates the payload environment producedby these sources during flight until the fairing isdeployed. After fairing deployment, the maximumfield strength produced by these sources at thepayload interface (Taurus station 0.0) is 10 V/min S-Band and 88 V/m in C-Band.

Ground Processing Environment — Specific RFenvironments experienced by the payload duringground processing at the PPF and the launch sitewill depend on the specific facilities that areutilized as well as operational details. Typicallythe field strengths experienced within the PPFduring payload processing and within the fairingat the launch site during final vehicle processingwill be less than 2 V/m from continuous sourcesand less than 10 V/m from pulse sources. Detailedenvironments for these operational phases willbe defined during the mission integration process.

RF SourceTx

PowerLocation

(x-station)

UG

34-1

C-Band TransponderS-Band TransmitterS-Band Transmitter* The C-Band transponder average power dissipation is 1.5 W

-28.7"-28.6"-28.6"

400 W (peak)*5 W5 W

Frequency5765 MHz2288.5 MHz2269.5 MHz

Figure 4-9. Taurus RF Sources.

Section 5Payload Interfaces


Taurus User's Guide Section 5.0—Payload Interfaces

5.1 Payload Interfaces

This section describes the available mechanical,electrical and LSE interfaces between the Tauruslaunch vehicle and the payload. Many of theseinterfaces are payload-specific and will bedetailed in the ICD developed for a given mission.

5.2 Payload Fairings

Taurus offers two payload fairing configurations,enabling the customer to optimize performanceand volume requirements. The 92” diameterpayload fairing provides the largest payloadenvelope in its class, while the 63” diameterfairing provides increased performance-to-orbitwith a smaller payload envelope. Both providesecurity and environmental control during groundprocessing, integration operations, and ascent.The fairings utilize graphite/epoxy compositeconstruction, are RF-opaque, and include internalacoustic blankets to control the payload acousticenvironment. The standard blankets are oneinch (25.4 mm) thick, but the thickness may betailored to meet mission-specific requirements asdiscussed in Section 8.3.3.

The two halves of the fairing are structurallyjoined along their longitudinal interface usingOrbital’s low contamination frangible jointsystem. An additional circumferential frangiblejoint at the base of the fairing supports the fairingloads. At separation, a gas pressurization systemis activated to pressurize the fairing deploymentthrusters. The fairing halves then rotate aboutexternal hinges that control the fairing deploymentto ensure that payload and launch vehicleclearances are maintained. All elements of thedeployment system have been demonstratedthrough test to comply with stringentcontamination requirements.

5.2.1 Payload Dynamic Envelope

Figure 5-1 and Figure 5-2 define the dynamicenvelope available to the payload in the 63” and92” fairings, respectively. The dynamic envelopesaccount for the deflection and manufacturingtolerances of the payload fairing and the payloadcone. The customer must verify that the payloadremains within the dynamic envelope when both

payload manufacturing tolerances and payloaddynamic deflections are taken into account.Coupled Loads Analysis results will be providedso that this verification can be performed.

Protrusions beyond the standard dynamicenvelope can be evaluated on a case-by-casebasis. Any payload extension aft of the payloadinterface also requires specific evaluation andagreement from Orbital.

5.2.2 Payload Access Door

One access door is provided for the payload as astandard service. For the 63” fairing, the size ofthis door is 12 inches x 12 inches (305 mm x 305mm). For the 92” fairing, the door is 18 inchesx 24 inches (457 mm x 610 mm), with the 18 inch(457 mm) dimension in the direction of thelaunch vehicle thrust axis and the 24 inch (610mm) dimension in the circumferential direction.The access doors are RF-opaque.

The standard door will be provided in thecylindrical section of the fairing, with the specificlocation determined through the MissionIntegration Working Group process. The locationof the door must be defined no later than L-15months.

Additional doors, doors of non-standard sizes ordoor locations other than in the fairing cylindercan be provided as non-standard services. Theseservices are discussed in Section 8.3.

5.3 Payload Mechanical Interface

Orbital will provide all hardware and integrationservices necessary to attach the payload to andseparate the payload from the Taurus vehicle.Orbital offers either a 38.81inch (986 mm) or37.15 inch (944 mm) diameter bolt circle payloadseparation system as a standard service. Orbitalcan also provide alternate size and capabilityseparation systems (23.25" and 38.81" highcapacity) and a 38.81" non-separating interfaceas optional services. These optional service aredetailed in Section 8.4. Payloads employingcustomer-furnished separation systems can alsobe accomodated, pending coordination of theinterfaces with the launch vehicle.



Figure 5-1. Payload Dynamic Envelope for 63" Diameter Fairing.UG-45

NegotiableAdditionalVolume

104.9(2664)

129.9(3299)

154.9(3934)

Sta 5.1Sta 0.0

ø30.3(ø770)

ø48.0(ø1219)

ø55.3(ø1405)

PayloadDynamicEnvelope

Fairing Envelope MotorPayload Separation System

Notes:

1. All Dimensions Are in Inches (mm)

2. All Station Numbers Are in Inches

3. Acoustic Blanket Thickness Is 1 Inch (25.4 mm)

4. Orbital Requires Definition of Payload FeaturesWithin 2 Inches (50.8 mm) of Payload Envelope

5. Projections of Payload Appendages Below Station 5.1May Be Permitted But Must Be Coordinated with the Taurus Program Office

6. One 12 Inch x 12 Inch (305 mm x 305 mm)Access Door Provided as Standard Service

Payload Envelope



Fairing Envelope MotorPayload Separation System

Notes:


2. All Station Numbers Are in Inches


4. Orbital Requires Definition of Payload Features Within 2 Inches (50.8 mm) of Payload Envelope

5. Projections of Payload Appendages Below Station 5.1May Be Permitted But Must Be Coordinated With theTaurus Program Office

6. One 18 Inch (457 mm) x 24 Inch (610 mm)Access Door Provided as Standard Service

Payload Envelope

Figure 5-2. Payload Dynamic Envelope for 92" Diameter Fairing.UG-21

Sta 5.1

ø19.2(488)

130.3(3310)

Sta .0.0

224.8(5710)

ø52.5(ø1334)

ø80.9(ø2055)

NegotiableAdditionalVolume

PayloadDynamicEnvelope

190.3(4834)



Orbital's 37.15" payload separation system isshown in Figure 5-3. The 37.15" and 38.81"standard configurations, which vary only in theirforward interface ring geometry, have the samestructural and separation performance. Theinterface flange geometries are provided in Figure5-4. The separation systems are manufacturedfor Orbital by Saab Ericsson Space of Sweden.Saab has extensive experience in supplyingseparation systems for a wide range of launchvehicles and payloads. The system is based on aseparation system that has flown over 30 timeswith 100% success.

The separation system is a marmon clamp banddesign that employs two aluminum interfacerings that are clamped by dual, semi-circularstainless steel clamp bands with aluminum clampshoes. Each of the two retention bolts is severedby a redundantly initiated bolt cutter. Uponband release, the movement and parking of eachband is controlled by a set of four extractionspring assemblies and two band catchers that aredesigned to prevent recontact with the upperring. Separation velocity is provided by up toeight matched spring actuators that impart up to27.7 ft-lb (37.6 Joules) of energy. The springassemblies may be tailored to mitigate the effectsof payload center of gravity (C.G.) offset. Thetypical tip-off performance of the systems is

described in Section 3.7. A typical layout of theseparation system is shown in Figure 5-5 whilethe structural capacity of the separation system isdetailed in Figure 5-6.

The mass of the separation system hardware thatremains with the payload is approximately 12.5lb (5.7 kg) for the 38.81" system and 10.0 lb (4.6kg) for the 37.15" system. This includes theinterface fasteners, upper interface ring, separationconnector bodies and connector bracketry.

UG41

Figure 5-3. Taurus Separation System.

37.80960

inmmDimensions

UG44

Figure 5-4. Interface Flanges for Taurus' StandardSeparation Systems.

Flange for 37.15" Separation System

39.451,002

38.81986

Flange for 38.81" Separation System

37.91963

36.10917

37.15944

φ

φ

φ

φ

φ φ

UG42

Figure 5-5. Typical Layout of Taurus SeparationSystem (38.81" System Shown with Forward RingPartially Removed to Show Lower Ring).

0

90 270

180

BoltCutters

(2X)

SeparationConnectorAssemblies

(2X)

PayloadSeparation

Springs(Up to 8X)

ClampBand

Catchers(4X)

Clamp BandExtraction Springs

(8X)



Interface fasteners may be inserted from eitherthe launch vehicle or payload side of the interface.The standard finish on the upper interface ring isAlodine 1200. It is also permissible to applythermal protection tape to certain non-criticalareas of the system. Other surface finishes areavailable on non-critical surfaces as a non-standard service.

Orbital will provide a drill tool to the payload forprecision drilling of the payload/launch vehicleinterface. Use of the drill tool is dependent oncoordination with other mission requirements.Delivery outside of the United States or to foreignpersonnel within the United States will be subjectto United States Department of State approval.

5.4 Electrical Interfaces

Orbital provides a flexible set of electrical interfacecapabilities to accommodate Taurus payloadrequirements. Figure 5-7 provides a graphicaldepiction of the electrical interface configurationand options. The physical interface to thespacecraft is provided by two separationconnectors. A mission-unique wire harness runsfrom the separation connectors to the Taurusavionics shelf, where signals are then distributedas appropriate to the T-0 umbilical connectors,the flight computer, and the telemetry multiplexer.

5.4.1 Payload Separation Connectors

As a standard service, Orbital provides two 56-pin low-impulse separation connector assemblies.The assemblies include a separation connectorwith strain relief backshell and all requiredmounting bracketry. Each connector provides48 20-gauge contacts and eight 16-gaugecontacts. The connector design is based upon astandard Mil-C-38999, 25-4 insert and uses atunable internal spring package to minimizeforces imparted during separation; thus, reducingpayload tip-off rates. Two of the 20-gaugecontacts in each connector are used by Orbital toverify payload separation via continuity loops,leaving up to 54 pins in each connector availablefor payload use.

Orbital will provide both halves of each separationconnector and all the mission-unique wiring onthe launch vehicle side of the interface. Theforward half of each flight connector will beprovided to the customer for integration into aninterface pig-tail harness. The aft side of eachflight connector will be integrated into the mission-unique launch vehicle wire harness provided byOrbital. If additional connectors are required aspart of the payload test program, they can beprovided as a non-standard service.

5.4.2 Umbilical Harness Pass-Throughs

The T-0 umbilical connectors provide the interfacebetween the launch vehicle wire harness and theground support harnesses to the LEV. Up to 64conductors (32 twisted shielded pairs), areavailable to pass payload signals to the payloadelectrical ground support equipment located inthe LEV. Four pairs of the 32 available wire pairshave a total round trip resistance of less thanthree ohms. These lines provide minimal powerloss to support battery charging, external power,and other current needs of less than four amps.The remaining 28 pairs have a round trip resistanceof less than 20 ohms. These lines are typicallyused for signal and communication.

5.4.3 Payload Command and Control

The Taurus flight computer can provide discretesequencing commands to the payload based on

Assumptions: Axial Accel = 2 g's (Tension), Lateral Accel = 3.5 g's, Load Peaking Factor = 1.0 (Where Load Peaking Is Defined as: Ratio of Maximum Predicted Line Load to Maximum Line Load Which Would Occur for an Ideal Uniform Structure.)

UG43

Figure 5-6. 38.81"/37.15" Separation SystemStructural Capability.

200(5.1)

150(3.1)

100(2.5)

50(1.3)

0500

(230)1,500(680)

1,000(450)

2,000(910)

Payload Mass, lb (kg)

Pay

load

C.G

., in

(m

)

2,500(1,140)



mission time or mission events. The discretecommands are optically isolated pulses ofprogrammable lengths in multiples of 40milliseconds. Up to eight optically isolateddiscrete outputs or switch closures are available,each capable of multiple pulses. The payloadsupplies the voltage (≤40 VDC) and the current

must be limited to less than 20 milliamps in afashion similar to using a dry contact relay.

5.4.4 Payload Discrete Status Monitoring

Up to eight optically isolated payload discretetelemetry inputs points can be monitored duringpre-launch operations and down-linked during

Launch Vehicle

5 Amp, 75 msec

5V

Spacecraft

* Optional for Certain Channels

Standard Quantity

8

8

8

8 Drivers

10Optional Non-Standard

Service

Separation Indicator

Pyro Initiation

Analog Telemetry0-5 V

-3.3 to 7.3 mV0 to 30 mV0 to 40 mV

-25 to 25 mV-79 to 162 mV

Serial Communication LinkRS-422/RS-485

GSE Pass-Through Wiring 64 Conductors(Various AWG/Shielding Available)

Discrete Commands

5VDiscrete Telemetry

Monitors (5 Hz)

1Optional Non-StandardCommunications Link

Figure 5-7. Taurus Electrical Interface Block Diagram.UG064

*



ascent. The eight discrete inputs can be ground-mode type, used for breakwire monitoring orvoltage-mode type used for bus "on/off"monitoring. In the ground-mode type, a 5.0 VDCsource is current limited to 500 microamps. Inthe voltage-mode type, the payload supplies theDC voltage and the current resistor to limit thecurrent to less than five milliamps. These discretesare typically monitored at five samples per secondand, depending on the launch vehicle telemetryformat, they can also be sampled at 25 samplesper second.

Note that critical telemetry points that must bemonitored at the customer’s convenience duringground processing should be routed through theumbilical pass-throughs, as the Taurus telemetryis available only during a limited number of pre-launch test operations.

5.4.5 Pyrotechnic Initiation Signals

Taurus has the ability to provide up to eightseparate initiation signals for electro-explosivetype devices. Up to four one-ohm electro-explosive type devices can be initiatedsimultaneously. The pulses are provided byTaurus' Pyrotechnic Driver Unit (PDU), which iscommanded by the flight computer through aserial communication link. The PDU is capableof delivering a pulse of 5 to 10 amps for 75milliseconds across a one-ohm load. This pulseis suitable for initiation of any standard electro-explosive device.

If the customer chooses to use the pyrotechnicinitiation service, the pyrotechnic signals mustbe segregated onto one of the two separationconnectors. If payload requirements do notpermit one of the two separation connectors tobe dedicated to the pyrotechnic signals, then athird separation connector can be provided as anon-standard service.

5.4.6 Payload Analog Telemetry Monitoring

As a non-standard service, up to ten analogtelemetry points may be monitored by thetelemetry multiplexer during pre-launchoperations and down-linked during ascent. The

multiplexer has eight-bit resolution for analogtelemetry and samples at five samples per second.The multiplexer has flexibility to accommodate arange of sensor types as detailed in Figure 5-7.Note that the ten available analog inputs consistof a combination of the various inputs and thatthe customer must provide the scale and biasvalues for the Taurus multiplexer.

Note that critical telemetry points that must bemonitored at the customer’s convenience duringground processing should be routed through theumbilical pass-throughs, as the Taurus telemetryis available only during a limited number of pre-launch test operations.

5.4.7 Serial Communication Interface

As a non-standard service, a serial RS-422/RS-485 communication interface between the Taurusflight computer and the payload can be provided.The flight computer will interrogate the payloadat a pre-determined rate and will receive payloadtelemetry to be interleaved into the downlinkedtelemetry stream.

5.4.8 Payload Battery Charging

Payload battery charging can be remotelycontrolled from the LSV. Payload chargingequipment, which is supplied by the customer,will be located in the LEV. The pass-throughlines from the LEV to the payload canaccommodate a maximum charging current offour amps. The pass-through conductorresistances are defined in Section 5.4.2.

5.4.9 Pre-Launch Electrical Constraints

Prior to launch, all payload and payload electricalground support equipment circuits shall beconstrained to ensure that the current flow acrossthe umbilical interface is less than or equal to tenmilliamps.

5.4.10 Pre-Separation Electrical Constraints

Prior to payload separation, all launch vehicleand payload interface circuits shall be constrainedso that the current across the separationconnectors is less than or equal to ten milliamps.



5.4.11 Electrical Launch Support Equipment

The Taurus electrical launch support equipmentconsists of two main operational units: the LSVand the LEV. Mission-unique, customer-suppliedpayload consoles and equipment can besupported in the LSV and LEV. Interface to thepayload through the Taurus T-0 umbilicals andland lines provides the capability for directmonitoring of payload functions. Orbital providesthe cabling from the T-0 umbilicals to the LEV,the junction box in the LEV, the fiber optic cablesfrom the LEV to the LSV, and the associatedcommunications and multiplexing equipment.Any additional mission-unique equipment mustbe provided by the customer.

Launch Support Van — The LSV houses all thehardware required to monitor and control thevehicle and payload.

The Payload Support Area (PSA) in the LSV iscapable of supporting two customer personnel.A removable 72 inch x 30 inch (1.8 m x 0.8 m)work surface is available for positioning thepayload-provided computer and monitoringequipment. Communication interfaces providedat the PSA consist of an RS-422 telemetry signal,an ethernet link, and up to 14 conductors of“patchable” signals which can be interfaced tothe LEV in a combination of RS-422 and TTL.Two intercoms are provided at the station. Theseintercoms can be set up per the customer'scommunication requirements. Power (120V/single phase) for payload equipment is availablevia several standard duplex receptacles.

Launch Equipment Van — The LEV houses theinterface between the LSV and the vehicle. Thisunmanned facility serves as a junction equipmentroom patching fiber optic lines from the LSV intoa copper connection with the vehicle. The LEVcontains a 60 in x 66 in x 32 in (1.5 m x 1.7 m x0.8 m){WxHxD} space for installation of payloadracks and associated equipment. Power—120V,single phase, 60Hz—is available via severalstandard 15 amp duplex receptacles.

The payload interface at the LEV is accomplishedthrough the patch panel and MUX/DeMUX. All

signals connected to the LSV will either use theMUX/DeMUX or the Optelcom Interface.

5.5 Payload Design Constraints

The following sections provide design constraintsto ensure payload compatibility with the Taurussystem.

5.5.1 Payload Center of Mass Constraints

The axial location of the payload center of massis typically constrained by the structural capabilityof the payload separation system. The separationsystem structural capability is defined as a functionof payload mass and center of gravity in Figure 5-6. If the Taurus payload separation system is notused, then the axial location of the payloadcenter of mass is constrained by the structuralcapability of the payload cone. The structuralcapability of the payload cone is defined inFigure 5-8.

The lateral offset of the payload center of mass isconstrained by the payload separation tip-off raterequirements and the structural capability of theseparation system. Because the payloadseparation system has a demonstrated capabilityof tailoring the separation springs to accommodatecenter of gravity offset and the additional momentloads due to center of gravity offset tend to besmall compared to the total capability of theseparation system, Orbital does not place a hard

Assumptions: Axial Accel = +5/-2 g's, Lateral Accel = 3.5 g's, Load Peaking Factor = 1.0 (Where Load Peaking Is Defined as: Ratio of Maximum Predicted Line Load to Maximum Line Load Which Would Occur for an Ideal Uniform Structure.)

UG066

Figure 5-8. Payload Cone Structural Capability.

250(6.4)

200(5.1)

150(3.8)

100(2.5)

50(1.3)

0500

(227)1,500(681)

1,000(454)

2,000(908)


Pay

load

C.G

., in

(m

)

2,500(1,135)



constraint on the lateral offset of the payloadcenter of gravity. If preliminary designassessments indicate that the lateral center ofgravity offset of the payload may exceed one inch(25.4 mm), the customer is encouraged to contactthe Taurus Program Office to verify the feasibilityof achieving the specific payload tip-offrequirements.

5.5.2 Final Mass Properties Accuracy

The final payload mass properties statement shallspecify the payload weight to an accuracy of0.5%, the center of gravity to an accuracy of 0.25inches (6.4 mm) in each axis, and the inertiamatrix components to an accuracy of 5%.

If the payload uses liquid propellant, the sloshfrequency shall be provided to an accuracy of 0.2Hz, along with a summary of the method used todetermine slosh frequency.

5.5.3 Payload EMI/EMC Constraints

The Taurus avionics RF susceptibility levels havebeen verified by test. The payload design mustincorporate inhibits that are at least single-faulttolerant to inadvertent RF radiation. The timeafter separation at which the payload may transmitwill be defined during the Mission IntegrationWorking Group process.

Prior to launch, Orbital will review the payloadradiated emissions (MIL-STD-461, RE02) to verifyoverall launch vehicle EMI safety margins inaccordance with MIL-E-6051.

Payload RF transmission frequencies must becoordinated with Orbital and range officials toensure non-interference with Taurus and otherrange transmissions. Additionally, the customermust schedule all RF tests at the processingfacility with Orbital in order to obtain properrange clearances and frequency protection. Oncethe payload is encapsulated within the payloadfairing, Orbital does not allow any payload RFtransmissions.

5.5.4 Payload Dynamic Frequencies

To avoid unfavorable dynamic coupling withlaunch vehicle forcing functions, the payload

should meet the dynamic frequency requirementsspecified in Section 4.4.

5.5.5 Payload-Supplied Separation Systems

Should the payload provide the separation system,the shock delivered to the payload cone interfaceflange must not exceed the levels defined inFigure 5-9.

5.5.6 System Safety Constraints

Orbital considers the safety of personnel andequipment to be of paramount importance. Eachcustomer is required to conduct at least onededicated payload safety review in addition tosubmitting to Orbital safety documentation assummarized in Section 7.7.3 and defined in EWR127-1.

Customers designing payloads that employhazardous subsystems are advised to contactOrbital early in the design process to verifycompliance with system safety standards.

EWR 127-1 outlines the safety design criteria forTaurus payloads. These are compliancedocuments and must be strictly followed. It is theresponsibility of the customer to ensure that thepayload meets all Orbital and range imposedsafety standards.

UG67

Figure 5-9. Maximum Payload Induced Flight Pyroshock Levels at the Payload Interface.

10,000

1,000

100

10100 1,000

Frequency (Hz)

SR

S L

evel

(g'

s)

10,000

Separable Interface Shock Levels

1003,500

10,000

SRS Level(Peak G)

Frequency(Hz)

5010,00010,000

Section 6Ground and Launch Operations


Taurus User's Guide Section 6.0—Ground and Launch Operations

6.1 Taurus/Payload Integration Overview

The Taurus system has been designed to minimizevehicle and payload handling complexity andlaunch base operations time. Horizontalintegration of the Taurus vehicle's upper stagessimplifies integration procedures, increases safetyand provides excellent access for the integrationteam. In addition, simple mechanical andelectrical interfaces and checkout proceduresreduce vehicle and payload integration times,increase system reliability and minimize vehicledemands on payload availability.

Orbital’s approach to integration also places fewrequirements on the payload. Payload processingis conducted a short distance away from thelaunch site in an environmentally controlledfacility such as Astrotech. Orbital encapsulatesthe payload and transports the assembly to thelaunch site 13 days before launch. A portable airconditioning unit supplies clean, temperature-and humidity- controlled air to the payload fromencapsulation through launch.

This section describes the integration facilities,the vehicle and payload integration activities,and the final integration operations performed atthe launch site in preparation for a Taurus mission.

6.2 Facilities

6.2.1 Western Range Facilities

The Western Range at VAFB is typically used forTaurus payloads requiring high inclination orbits.Taurus launches from the Western Range areconducted from Space Launch Complex 576E(SLC 576E), located on the coast of North VAFB(NVAFB). Payload processing for commerciallaunches is normally performed at AstrotechVAFB, a commercial payload processing facility.The Missile Assembly Building (MAB), locatedon NVAFB, is used for Taurus stage integrationactivities.

6.2.1.1 Payload Processing

A commercial facility, Astrotech SpaceOperations VAFB (ASO), is the primary Payload

Processing Facility (PPF) for commercial Tauruslaunches conducted from the Western Range.Astrotech VAFB is located near the northwestend of the VAFB airfield at the corner of Tangairand Red Roads, approximately 3.5 miles fromSLC 576E. This facility has been successfullyused for processing payloads launched on Taurus,Pegasus, Delta and Atlas launch vehicles.

The Astrotech complex consists of the followingmajor structures:

a. Payload Processing Facility: Used for allpayload preparation operations, hazardousoperations, payload fueling, etc.

b. Technical Support/Customer Office: Locatedin the non-hazardous work area, this facilitysupports customer administrative functions,as well as serving as ASO administrativeheadquarters.

c. Warehouse: Used for ASO storage, but maybe used for customer GSE storage.

The PPF, depicted in Figure 6-1 and detailed inFigure 6-2, consists of the following customerareas:

a. West Highbay: The West Highbay is a 2400ft2 (223 m2) cleanroom with an adjacentdedicated control room. Two cleanroomlowbays are attached to West Highbay, eachwith highbay access through 13.5 x 10 ft (4.1x 3.0 m) roll-up doors.

b. East Highbay: The East Highbay is a 3500 ft2

(325 m2) cleanroom with an adjacentdedicated control room. A single cleanroomlowbay is attached with highbay accessthrough a 13.5 x 10 ft (4.1 x 3.0 m) roll-updoor.

c. Airlock: The airlock is a 2400 ft2 (223 m2)cleanroom capable area with a 20 x 40 ft (6.1x 12.1 m) exterior roll-up door and a 20 x 45ft (6.1 x 13.7 m) roll-up door to the WestHighbay.

d. Change Rooms/Air Showers: Each highbayhas a dedicated change room with lockersand an air shower that can be used prior tohighbay entry.



UG054

Figure 6-1. Astrotech VAFB Payload Processing Facility.

East High Bay3500 sq ft(325 sq m)

West High Bay2400 sq ft(223 sq m)

MechanicalRoom

400 sq ft(37 sq m)

Control Rm1000 sq ft(93 sq m)

Change RoomsAir Showers

Control Rm650 sq ft(60 sq m)

Mechan-ical

Room

Elec.Room

Auxiliary Control Room2160 sq ft(200 sq m)

SharedAirlock

2400 sq ft(223 sq m)

Low Bay

Low Bay400 sq ft(37 sq m)

400 sq ft(37 sq m)



6.2.1.1.1 Security

ASO has a Defense Investigative Service facilityclearance designation. Access to the Astrotechcomplex and the PPF is controlled through a cardreader access system. The customer can specifythe individuals who are authorized access totheir specific control room and clean room areasduring facility usage.

6.2.1.1.2 Thermal and Cleanliness

The air conditioning system is computercontrolled and monitored and maintains thecleanroom areas to 70 ± 2° F (21 ±1°C) and 45 ±10 % relative humidity. A positive pressure ismaintained in all clean room areas with aircirculated through HEPA filters at a rate of 3.5 –4.0 air changes per hour.

Cleanliness in the bays and airlock is certified toClass 100,000 and is functional to below Class10,000.

6.2.1.1.3 Cranes

The airlock and West Highbay are equipped witha 10-ton (9,100 kg) overhead trolley bridge cranewith a 37 ft (11.3 m) hook height. The EastHighbay is equipped with a 30-ton (27,000 kg)overhead trolley bridge crane with a 55 ft (16.8m) hook height. Both cranes have multi-speedcontrol including microspeed control. Forpayload personnel designated to perform craneoperations, Astrotech will conduct craneoperation training classes.

6.2.1.1.4 Electrical

Astrotech is served by a 480 Volt/3 phase facilitypower feed. Figure 6-3 indicates isolated andsurge protected power sources that are availablein the PPF to accommodate payload needs. ThePPF is supported by a 250 kW diesel poweredbackup generator.

Floor Area ft Fueling Island ft

Cleanliness

Temperature °F °C

Relative Humidity

CraneCapacity/Hook Height

Exterior Doors ft

Interior Doors ft

WestHighbay40 x 60

12.1 x 18.325 x 25

7.6 x 7.6100,000 *

70 ±221 ±1

45 ±10%

10 ton/37 ft(9,100 kg/11.3 m)

N/A

13.5 x 104.1 x 3.0(lowbays)20 x 45

6.1 x 13.7(airlock)

50 x 7015.2 x 21.3

25 x 257.6 x 7.6100,000 *

70 ±221 ±1

45 ±10%

30 ton/55 ft(27,000 kg/

16.8 m)

20 x 506.1 x 15.2

20 x 456.1 x 13.7

(E. Highbay)13.5 x 104.1 x 3.0(lowbay)

40 x 6012.1 x 18.3

25 x 257.6 x 7.6100,000 *

70 ±221 ±1

45 ±10%

10 ton/37 ft(9,100 kg/11.3 m)

20 x 406.1 x 12.1

See E. Highbay

20 x 236.1 x 7.0

6 x 61.8 x 1.8100,000 *

70 ±221 ±1

45 ±10%

N/A

8 x 102.4 x 3.0

24 x 277.3 x 8.2

N/A

N/A

50-75 ±210-24 ±1

**

N/A

8 x 102.4 x 3.0

23 x 42.57.0 x 13.0

N/A

N/A

50-75 ±210-24 ±1

**

N/A

8 x 102.4 x 3.0

EastHighbay Airlock Lowbays

WestControl Room

EastControl Room

* Rooms are functional to below class 10,000 ** RH control capability is dependent on the selected temperatureUG010

Figure 6-2. Astrotech VAFB PPF Details.

m

m

m

m

( )( )

( )

( )

( )



6.2.1.1.5 Communications

Telephone/Fax — Phones for voice and facsimileare available throughout the ASO complex. Thecustomer can set up voice mail systems if desired.

Voice Net — ASO maintains an operationalvoice net system capable of similtaneous accessto 12 voice nets that may connect to variousWestern Range facilities.

Paging — A paging system is available andoperable throughout the complex.

Closed Circuit Television — Two color cameraswith remote control zoom lenses and pancapability are installed in each highbaycleanroom. A single camera is installed in theairlock. The video signals can be transmitted toany building in the ASO complex.

Fiber Link — ASO is connected to the VAFB fiberoptic cable communications network and datacan be transmitted to any end use location on thenetwork.

6.2.1.1.6 Payload Fueling Capability

The ASO facility is fully capable of supportingpayload fueling operations. As a non-standardservice (see Section 8.9), ASO provides a turn-key service for customers to support propellantloading operations to include the following:propellant transport, sample transport andanalysis, life support services, loading facilities,and decontamination/waste handling services.

For those payloads requiring it, hydrazine isavailable from ASO through their agreementwith the Air Force. Propellant costs are passeddirectly to the customer with no additionalcharges.

6.2.1.1.7 Other Equipment and Services

Other equipment and services provided by ASOinclude:

• Forklift – 3,000 lb (1,360 kg) capacity,cleanroom compatible;

• Highreach – electrically driven andcleanroom compatible;

• Compressed Air – for breathing, shop, andpallet jack operations; and

• Compressed Gases – limited quantitiesmeeting applicable MIL specifications areavailable.

6.2.1.2 Stage Integration Facility

The stage integration facility for all Tauruslaunches is the MAB located on NVAFB andshown in Figure 6-4. Taurus uses this facility tohorizontally integrate and test the launch vehiclecomponents prior to their delivery to the launchsite for final launch operations. The building is356 ft (109 m) long by 138 ft (42 m) wide with apeak height of 143 ft (43.6 m). There are twooverhead bridge cranes with hook heights of 104ft (31.7 m) and capacities of 50 tons (45.4k kg)and 150 tons (136k kg).

The facility houses an array of Taurus GSE,including:

• An Assembly and Integration Trailer (AIT),stationary rails,and motor dollies forprocessing of Taurus hardware;

• Equipment for transportation, delivery,loading, and unloading of the Taurus vehiclecomponents;

• Equipment for integration and test of a Taurusvehicle; and

• Equipment to maintain standard payloadenvironmental control requirements.

6.2.1.3 Launch Site Facilities

SLC 576E, shown in Figure 6-5, is the primarylaunch facility for Taurus at the Western Range.This facility, formerly used for launching AtlasICBMs, is relatively austere with few permanentstructures. SLC 576E contains a launch padsurrounded by security fencing, lighting andcamera towers, and is supported by facility powerand the VAFB Communications Network.

110 Volt/1 Phase/60 Hz208 Volt/2 Phase/60 Hz480 Volt/3 Phase/60 Hz

Standard 20 amp Duplex Receptacles

Crouse Hinds BHRE 658NW, 60 Amp

Crouse Hinds BHRE 3583NW, 30 Amp

Power Source Receptacle

UG011

Figure 6-3. Astrotech PPF Power Sources.



6.2.1.3.1 Launch Support Equipment

During launch campaigns, Orbital supplies andinstalls mobile Launch Support Equipment (LSE)to provide all the additional site support necessaryfor launch operations. The primary LSEcomponents are summarized in the followingparagraphs.

LEV — As previously described in Section 2, theLEV contains both launch vehicle and payloadinterface equipment and serves as the fiber-to-copper interface between the LSV and the launchvehicle. The LEV is positioned approximately200 ft (61 m) from the launch stand and isunmanned during launch operations.

Launch Stand — The Taurus launch stand isinstalled on the launch pad well before flight

hardware arrives at the site.

Integration Tent — This clamshell tent serves asthe facility for final vehicle testing and payload-to-vehicle mating operations. The tent is locatedadjacent to the launch stand and is moved priorto upper stack mating.

Scaffolding — An OSHA approved scaffold isconstructed around the Stage 0 motor and 0/1interstage, providing pre-launch access to thesestructures. The scaffold is removed prior tolaunch.

Aerial Manlift Trucks — These vehicles are usedto access the Taurus upper stages and fairingareas once the entire launch vehicle has beenvertically erected on the launch stand (typicallyat L-6 days).

UG090

Figure 6-4. Missile Assembly Building.



Equipment Support Trailer — Orbital provides aportable container for housing tools and otherLSE.

Administrative Trailer — Located outside thehazard operations zone, this facility providesadministrative space for Orbital and customerrepresentatives. This trailer contains telephone,facsimile, computer, restroom and otherresources.

6.2.1.3.2 Electrical

SLC 576E has a 60 Hz facility power feed whichis distributed throughout the launch complex at208-480V/3phase and 110-208V/1phase. Allflight critical equipment/components (e.g., LEV,Payload Environmental Control System, etc.) aresupported on an emergency backup powergenerator system which automatically transferspower within seconds of sensing a loss of facilitypower.

Specific payload GSE power requirements shouldbe coordinated with Orbital early in the missionplanning process.

6.2.1.3.3 Communications

Telephone/Fax: Telephone and facsimile supportis available at the Administrative Trailer.

LSV Launch Control: Launch vehicle and payloadcommunications from the LSV are providedthrough the LSV-LEV fiber optic cable and LEV-launch vehicle umbilical cables.

Fiber Optic Network: SLC 576E contains atermination point for the VAFB fiber opticcommunications network.

6.2.1.3.4 Security

The launch site is surrounded by a chain linkperimeter fence with a 3-strand barbed wire top.During periods when flight hardware is at thesite, access is controlled by a security force.

CameraTower

LEV

Crane

AdminTrailer

BackupGenerator

IntegrationTent

Launch Standand Scaffolding Abandoned

Missle Silo UG099

Figure 6-5. SLC-576: Primary Launch Facility for Taurus at the Western Range.



6.2.2 Eastern Range Facilities

Taurus integration operations at the Eastern Rangemirror those conducted at the Western Range.The obvious difference is that payload processingas well as the final vehicle integration and launchoccurs at CCAS.

For an Eastern Range launch, Orbital integratesthe Taurus stages at the MAB on VAFB. Once thevehicle has completed stage integration andpreliminary systems testing, the integrated Taurusstages are delivered to the launch site on CCASand the Taurus payload cone and fairing aredelivered to the PPF at CCAS.

Like Astrotech on VAFB, Orbital will provide aPPF at CCAS for commercial Taurus customers.The PPF will provide the same basic support andservices (described in Section 6.2.1.1) requiredto integrate, test and encapsulate the payload.

For CCAS launches, Orbital will use thecommercial Launch Complex 46 (LC-46),developed by the Spaceport Florida Authority inconjunction with launch industry partners(including Orbital). LC-46, shown in Figure 6-6and Figure 6-7, is a proven commercial launchfacility capable of supporting a variety of launchvehicles. Orbital has signed an agreement withSpaceport Florida to use LC-46 for Taurus missionsthat require an East Coast launch. Followingintegration at the MAB, Taurus hardware will beshipped to CCAS via rail and truck. Once at LC-46, Orbital will perform standard launch siteintegration procedures to verify the hardware

was not damaged during transport, mate theencapsulated cargo element with the Taurusvehicle, conduct the integrated systems tests,and perform the final erection and close-out ofthe vehicle.

6.3 Vehicle and Payload Integration Operations

Taurus ground and launch operations areconducted in three major phases:

• Launch Vehicle Integration: Assembly andtest of the Taurus vehicle at the MAB;

• Payload Processing: Receipt and checkoutof the payload at the PPF followed byencapsulation into the Taurus fairing; and

• Launch Operations: Transport of vehiclecomponents and the encapsulated payloadto the launch site, final integrated systemstesting, erection of the vehicle onto the launchstand, and launch.

Each of these phases is more fully described inthe following sections. Orbital maintains launch

UG110

Figure 6-6. LC-46 Launch Complex.

UG111

Figure 6-7. LC-46 Tower.



site management and test schedulingresponsibilities throughout the entire integrationcycle. Additionally, all Orbital integrationactivities are controlled by a comprehensive setof Taurus Work Packages (TWPs), which describeand document every aspect of integrating Taurusand its payload. Mission specific work packagesare developed for mission unique or payloadspecific operations.

6.3.1 Launch Vehicle Integration

All launch vehicle motors, parts, and completedsubassemblies are delivered to the MAB fromeither the assembly vendor or Orbital’s Chandlerproduction facility. Figure 6-8 depicts the typicalflow of hardware from the factory to the launchsite.

The transformation of solid rocket motors,avionics, and sub-assembled structures into anintegrated launch vehicle begins at the MAB.Figure 6-9 outlines the typical activity flow forMAB operations while Figure 6-10 depicts atypical vehicle configuration during MAB testing.A small group of skilled engineers and techniciansperform the following major functions:

• Receive and inspect all motors, sub-assemblies and vehicle components;

• Integrate mechanical and electricalcomponents and sub-assemblies onto theindividual motors;

• Perform electrical and systems testing of theintegrated motors, composite sub-assembliesand the avionics section; and

• Physically and electrically mate the vehicleupper stages (i.e., Stages 1, 2, 3, and theavionics section).

The Taurus upper stages are integratedhorizontally at the MAB prior to their shipment tothe launch site. This integration is performed ata convenient working height which allows foreasy access for component installation, test andinspection. The integration and test processensures that all vehicle components and sub-systems are thoroughly tested before and afterfinal flight connections are made.

Vehicle systems tests verify operation of allsubsystems prior to stage mate. For each of thesetests, a specialized test software load is installedinto the Flight Computer. The major tests include:

• Vehicle Verification: A test that efficientlycommands all subsystems (TVCs, RCS, pyrocommands, etc.) in an accelerated timeline.

• Phasing Test: A test to verify the sign of thecontrol loop of the flight actuators anddynamic operation of the IMU. In this test theIMU is moved manually while the motion ofthe flight actuators (TVCs and RCS) is observedand recorded.

• Flight Simulation Testing: This series of testsuses the actual flight code and simulates a“fly to orbit” scenario. All flight actuators,pyro commands and flight computercommands are exercised. A Flight Simulationis repeated after each major vehicleconfiguration change (i.e., Flight Simulation#1 occurs after subsystem and componentinstallation; Flight Simulation #2a occursafter upper stage mate; Flight Simulation #2boccurs after transport to the launch site; andFlight Simulation #3 occurs after the payloadis electrically mated to the vehicle). Aftereach test the configuration of the vehicleremains unchanged until a full and competedata review of the test is complete. Thepayload nominally participates in FlightSimulation #3.

The successful completion of Flight Simulation#2a completes MAB operations. At this point,the fairing and payload cone assemblies areshipped to the PPF for payload encapsulationoperations. The Stage 0 motor, the Stage 0/1interstage, and the upper stage assembly areprepared for shipment to the launch site.

6.3.2 Payload Processing Operations

The integrated payload processing activities,outlined in Figure 6-11, are designed to simplifyfinal launch processing while providing acomprehensive verification of the payloadinterface. Payload processing is conducted



Fairing

Fairing

ThiokolAlliantOrbital

Stages1, 2 & 3

Stage 0Avionics

SpacecraftAdapter

SpacecraftAdapter

Stage Integrationand Systems Testing

SpacecraftOrdnance

Assemble, Integrate,Test, Install Ordnance,

Fuel, Encapsulate

PayloadProcessingFacility

EncapsulatedCargo

Element

Spacecraft

Contractor

Match-Mate TestInterface Verification

TaurusLaunch Site

Flight SpacecraftCheckout & HandlingEquipment

InterfaceMaster

Drill Tool

MissileAssemblyBuilding

UG080

Figure 6-8. Hardware Flow - Factory to Launch Site.



independently of Taurus vehicle processing,allowing the payload to determine and controlthe length of time required for checkout withinthe PPF. Once the payload has completed itsown independent verification and checkout,Orbital will deliver the fairing and payload coneassemblies to the PPF. The nominal PPF integratedoperations include mating the payload to theTaurus payload cone, performing interfaceverifications, and encapsulating the payloadwithin the Taurus fairing halves. Orbital willthen transport the encapsulated payload to thelaunch site. The standard launch service includesup to 30 days in the PPF with four of these daysslated for payload-to-Taurus integrated activities.

For planning purposes, typical PPF activities andtheir timelines follow.

• L-42 days: Payload arrival, preparation, andcheckout

• L-20 days: Fairing and payload cone deliveryto the PPF

• L-16 days: Begin payload mate andencapsulation activities

• L-13 days: Encapsulation complete; readyfor transport to launch site.

Payload Integration and Encapsulation — AfterFlight Simulation #2a is completed at the MAB,the fairing and payload cone assemblies arereleased and delivered to the PPF with theassociated GSE. In parallel with other PPFoperations, the fairing halves are lifted from theiras-shipped horizontal configuration to verticaland secured in Orbital-provided handling carts.Orbital then performs fairing system checks andstores the fairing halves until needed forinstallation as shown in Figure 6-12. The payloadcone is strategically placed in the PPF to facilitateboth payload mate and fairing installation. Oncethe payload cone assembly has been positioned,and the separation system has been installed, thepayload is lifted from its GSE stand (using customerprovided lifting fixtures), and mated to the payloadcone assembly, as shown in Figure 6-13.

Following the payload-to-Taurus mechanical andelectrical mating and interface verification testing,the payload is ready for encapsulation within thefairing. The first fairing half is rolled into placeand mated, as depicted in Figure 6-14. Followingany final payload closeout operations, the secondfairing half is rolled into place and mated,

UG050

Figure 6-9. Vehicle Processing Flow at the MAB.

L-4 Months L-2 Months

Stage Integration

Pre-SystemsTest Setup

StageTelemetry

Tests

FlightComputerSoftware

Load

RF and FTSSystems

Test

FlightSimulation

#1

Mate Stages1, 2 & 3

FlightSimulation

#2a

VehicleVerification

Test

Receive Stages& Hardware at

MAB

L-75 Days

L-20 Days

Receive Fairingand Payload

ConeAssemblies

Ship Stage 0, Stage 0/1 Interstage & Integrated

Upper Stages toLaunch Site

L-20 Days

Ship Fairingand PayloadCone to PPF

Flight ControlPhasing Test



completing the encapsulation process, as shownin Figure 6-15. Final assembly of the handlingfixture completes the Encapsulated Cargo Element(ECE). At this point, access to the payload isavailable only through the fairing access door.

The ECE may remain at the PPF in thisconfiguration for any length of time until transportto the launch site is required.

6.3.3 Payload Transport to the Launch Site

In preparation for transporting the ECE, Orbitalverifies that the Taurus mobile environmentalcontrol system (ECS) meets applicabletemperature, humidity, and cleanlinessrequirements. The system is then attached to thefairing via insulated flexible ducting and beginsproviding environmental control to the payload

UG089

Figure 6-10. Stage Integration in the MAB.

UG051

Figure 6-11. Payload and LV Processing Flow at the PPF.

ReceiveFairing/

Payload Cone

Place Fairingin Vertical

Carts

FairingDeployment

Checks

L-20Days L-13 DaysL-14 Days

Ship Payloadto Launch

Site

L-15Days

Fairing Installation/

Encapsulation

L-16 Days

PayloadArrival at

PPF

PayloadCheckout and

Testing atPPF

Mate Payloadto Payload

Cone

L-42 DaysPayload/Taurus

Interface Verification

Checks

PrepareFor Shipment

to LaunchSite



volume. The payload volume temperature andhumidity levels will be monitored and controlledin this manner for the remainder of groundoperations. An Orbital provided air-ride trailer,already loaded with the ECS and power generator,is then backed into the PPF airlock. The ECE ishoisted onto the trailer, secured, and preparedfor road transportation to the launch site, asshown in Figure 6-16. A limited amount of spaceis available on the transporter for any requiredpayload monitoring GSE. Power (120V/1phase/60hz) is also available for powering GSE duringthe transport.

6.4 Launch Operations

Launch operations consist of the following majoractivities:

• Verification of LSE, facilities, and utilities;• Delivery of the Taurus launch vehicle

components to the launch site for final

integration and test;• Transportation of the ECE to the launch site

for testing and integration with the Taurusvehicle;

• Final system testing and vehicle mate; and• Vehicle arming and launch.

6.4.1 Facility and LSE Checkout

Prior to the arrival of any flight hardware at thelaunch site, Orbital erects the launch stand andintegration tent, verifies the power and electricalinterfaces, and installs the remaining LSE. Orbitalalso thoroughly tests communications betweenthe Range, LSV, LEV and SLC 576E. Payload GSEinterfaces with the LSV and LEV must also beverified well in advance of flight hardware deliveryto the site. Orbital recommends that theseinterface verifications occur prior to the payloadarrival at the PPF.

6.4.2 Delivery of Taurus to the Launch Site

The launch site processing flow is detailed in

UG081

Figure 6-12. Fairing Stowed in PPF.

UG082

Figure 6-13. Payload Mated to Payload Cone.



Figure 6-17. Within three weeks of launch, theintegrated Taurus hardware is transported fromthe MAB to the launch site in three major pieces:the Stage 0 motor, the Stage 0/1 interstage andthe upper stage assembly. Upon arrival at thesite, the Stage 0 motor and Stage 0/1 interstageare immediately erected onto the launch standand the upper stage assembly is positioned insidethe integration tent in preparation for final vehicletesting. At this time, Stage 0 is electricallyconnected to the upper stages using a jumperharness and the T-0 umbilicals are connected,establishing vehicle control from the LSV. Orbitalthen performs a flight simulation test (FlightSimulation #2b) to verify that the vehicle healthwas unaffected by the transport. The test alsoverifies communication with and control of thevehicle from the LSV. In addition, Orbital verifiesthat the interfaces of the payload GSE meet theICD requirements. These tests, customized foreach mission, typically checkout the LSV/LEVcontrols, launch vehicle sequencing, and anyoff-nominal modes of the payload. It is

recommended that the customer complete thisactivity for the payload side of the interface aswell.

6.4.3 Payload Arrival and Activities at theLaunch Site

Once Flight Simulation #2b is successfullycompleted, Taurus is “ready” for payloadtransport. This typically occurs at L-13 days. Aspreviously described, the ECE is loaded verticallyonto its transporter at the PPF and is then

UG084

Figure 6-15. Payload Encapsulation Complete.

Figure 6-16. ECE Transport to Launch Site.UG085

UG083

Figure 6-14. First Fairing Half Installation.



transported to the launch site via convoy. Onceat the site, the ECE is crane-lifted from thetransporter, rotated to a horizontal position, andplaced next to the Taurus Upper Stage Assemblyinside the Integration Tent, as depicted in Figure6-18. Flight electrical connections are thenmade between the ECE and the launch vehicle.Once the ECE is electrically mated to Taurus,battery charging and communication with thepayload are nominally controlled from thepayload station in the LSV.

6.4.4 Integrated Testing and Vehicle Mate

A final integrated systems test (Flight Simulation#3) is performed with the vehicle in a full flightconfiguration (e.g., flight cables mated, systemson internal power, RCS firing, etc.), allowing acomplete check of all interfaces, including thepayload-to-launch vehicle interface. The payloadparticipates in Flight Simulation #3 by being in apowered mode and monitoring telemetry.Following the successful completion of this test,the ECE is mechanically mated to the Taurusupper stage assembly.

Using a special handling fixture and two cranes,

the upper vehicle stack with ECE is hoisted,rotated to vertical and mated to the previouslyerected Stage 0, as shown in Figure 6-19 . Payloadenvironmental control and electrical continuityis maintained throughout the hoisting and matingoperations. Final end-to-end continuity andfunctional tests are performed following themating.

6.4.5 Vehicle Arming

Following final vehicle testing, Taurus is armedand the pad area is cleared for launch. These

L-20 Days

L-4 Days L-3 Days L-2 Days L-1 Day L-0 Day

L-11 Days L-9 Days L-8 Days L-7 Days L-6 Days

L-19 Days L-13 Days

Payload Vertical

ErectStage 0

Install 0/1Interstage

Pre-SystemsTest Setup

FlightSimulation

#2b

L-12 Days

ElectricallyMate Payload

toTaurus Vehicle

FlightSimulation #3

Top OffExpendables(Batteries &

RCS)

MechanicallyMate

Payloadto Taurus

OrdnanceConnectionand Vehicle

Closeout

Mate UpperStack withStage 0

Receive 0/1Interstage atLaunch Site

ReceiveUpper-Stack atthe Launch Site

Receive Payloadat Launch Site

L-20 Days

Receive Stage 0at Launch Site

L-12 Days

PayloadInterface

VerificationTests

Final RF andFTS Tests

MissionRehearsal

LaunchReadiness

Review/Final Preps

Final Vehicle/PayloadArming

Launch0/1 Electrical

Tests

Rotate Payloadto Horizontal

UG049

Figure 6-17. Launch Site Processing Flow.

Figure 6-18. ECE in Position for Integration Withthe Launch Vehicle.

UG086



activities, which typically commence on L-1 dayand conclude by L-12 hours, take the Taurusvehicle to a launch-ready configuration as shownin Figure 6-20. This timeframe also representsthe final period for payload access under normalsituations.

The launch countdown is performed from theLSV located outside the caution/hazard corridor.

6.4.6 Launch

Following Taurus and payload closeout activitiesduring L-1 day operations, the final countdownprocedure begins. Figure 6-21 lists the keyTaurus tasks completed during the final two hourcountdown. These activities methodicallytransition the vehicle from a safe state to that oflaunch readiness. Payload tasks, as necessary,are included in the countdown procedure andare coordinated by the Taurus Test Conductor.

6.4.6.1 Launch Control Organization

The Launch Control Organization consists of twogroups with distinct responsibilities. Figure 6-22shows the typical communication flow requiredto make the launch decision.

The management group includes key personnelrepresenting the payload, launch vehicle, andrange. This group, headed either by the customerMission Director (MD) or the Orbital LaunchDirector, has the final decision authority for thelaunch. The management lead polls the necessarypayload personnel to ascertain the status of thepayload and receives the final Go/No-Gorecommendation from the Payload LaunchDirector. The appropriate launch rangeCommander provides the final rangerecommendation for launch. Thisrecommendation encompasses all aspects of therange’s responsibility including, but not limited

UG088

Figure 6-19. Upper Stack Lift.



to, flight safety, ground safety, hazard corridorstatus, local and downrange telemetry, trackingand command assets, and weather. The OrbitalLaunch Director provides the launch vehiclestatus and launch recommendation based oninputs from the Taurus launch team.

The second group is the operations/engineeringstaff, including the Test Conductor, the Taurusengineering team, payload engineering team,and the Range Control Officers. The TestConductor, an Orbital employee, is responsiblefor running the countdown procedure. The OrbitalVehicle Engineer has the overall responsibilityfor Taurus readiness while the Taurus OperationsEngineer performs the necessary commands toprepare the vehicle for launch. The Taurus TestConductor, Vehicle Engineer, and OperationsEngineer reside in the LSV. In addition, both theTaurus and Payload Launch Directors arepositioned in the LSV. There is also space for upto two additional payload support team membersin the LSV. These payload personnel provide

satellite status and readiness to the PayloadLaunch Director.

-120 min.

-90 min.

-75 min.-45 min.-35 min.-26 min.

-15 min.-10 min.-8 min.-7 min.-5 min.-4 min.

-2 min.-1 min.

T0

Countdown BeginsPrelaunch Operation Checklist Verified CompleteCall to StationsFinal Site Clear and LEV ActivationVehicle on External Avionics PowerRange Evaluation of TelemetryC-Band Transponder CheckoutWeather BriefingVerify Upper Level Winds Acceptable for LaunchFTS to Internal PowerFTS-Open Loop Check on Internal PowerEnable LSV for LaunchC-Band Interrogation Through LaunchAvionics Bus to Internal PowerFinal Range TLM EvaluationFinal Launch Go/No-Go PollTransient Power OnVehicle S&As ArmedAuto Sequence InitiatedIMU to Free InertialLiftoff

NominalTime Major Events

UG053

Figure 6-21. Final Countdown Sequence.

ECS

UG087

Figure 6-20. Taurus Ready for Launch.



The remainder of the launch team is positionedin the Range Control Center. Figure 6-23 showsa typical Range Control Center configuration fora VAFB launch. The Mission Director consoleand (four) additional customer consoles aretypically available for customer use. Additionalspace for key customer personnel and dignitariesis available in the Technical Support Room.

6.4.6.2 Launch Rehearsals

Two rehearsals are conducted prior to eachlaunch. The first is conducted at approximatelyL-1 month and is used to acquaint the launchteam with communications systems, reporting,problem solving, launch procedures andconstraints, and the decision making process.The rehearsal is a full day in duration and consistsof a number of countdowns performed usingabbreviated timelines. All aspects of the team’sperformance are exercised, as well as hold, scrub,and recycle procedures. The operations arecritiqued and lessons learned are incorporated

Vehicle Engineer

Mechanical Operations LaunchWeatherOfficer

Missile FlightControl Officer

Senior MissileFlight Control

Officer

FlightSafetyProject Officer

OperationsSupportManager

RangeControlOfficerPropulsion Electrical Systems

Located in LSV

Go/NoGo Coordination UG070

Figure 6-22. Typical Launch Decision Flow.

Customer MissionDirector

Orbital LaunchDirector

Test Conductor

Payload LaunchDirector

Payload SubsystemEngineers

Range OperationsCommander

RangeCommander

GN&C Flight Safety

prior to the Mission Dress Rehearsal (MDR) at L-2 days. The MDR is the final rehearsal prior toentering the countdown and will ensure problemsencountered during the first rehearsal have beenresolved. The MDR is typically a half day induration. All customer personnel involved withlaunch day activities participate in both rehearsals.



UG071

Figure 6-23. Typical VAFB RCC Seating Arrangement.

DISPLAYS DISPLAYS

Orbital

ENVIRONMENTS

STATION 19

TECHNICALSUPPORTROOM

(18 SEATS,COMM,VIDEO)STATION 29A

Orbital

STATION 26

Mission

Director

STATION 25

Orbital

RANGE

INTERFACE

STATION 17

Customer

STATION 16

Orbital

THERMAL

STATION 18

Orbital

WINDS-1

STATION 20

Orbital

WINDS-2

STATION 21

Orbital

STATION 22

Orbital

STATION 23

Orbital

STATION 24

Customer

STATION 13

Customer

STATION 15

Customer

STATION 14

STATION 27

OrbitalGNC-1

STATION 04

OrbitalPROPULSION

STATION 06

OrbitalGNC-2

STATION 05

OrbitalSYSTEMS

STATION 07

OrbitalPOWER

STATION 08

OrbitalELECTRICAL

STATION 09

OrbitalLEADELECTRICAL

STATION 10

OrbitalCABLES

STATION 11

OrbitalTVA

STATION 12

OrbitalMECHANICAL

STATION 01

OrbitalCHIEFENGINEERSTATION 02

OrbitalFLIGHTASSURANCESTATION 03

Section 7Mission Integration


Taurus User's Guide Section 7.0—Mission Integration

7.1 Management Approach

The Taurus program is managed through theimplementation of an integrated project plan.This plan is prepared by the Program Manager,based on inputs from the program team, andapproved by the General Manager of Orbital’sLaunch Systems Group. This project plan includesmaster schedules, the Work Breakdown Structure(WBS) dictionary, and organization and staffingplans and provides the baseline for trackingprogress in the areas of technical and scheduleperformance.

7.2 Orbital Mission Responsibilities

A mission unique organizational structure isestablished for each Taurus mission to manageand execute key mission roles and responsibilities.Open communication between Orbital and thecustomer, emphasizing timely transfer of dataand prudent decision-making, ensures efficientlaunch vehicle/payload integration operations.As the launch service provider, Orbital’sresponsibilities fall into four primary areas:

1) Program Management;2) Mission Management;3) Engineering; and4) Launch Site Operations.

The Taurus program organization is depicted inFigure 7-1, while the four areas of responsibilitiesare summarized in Figure 7-2 and in the followingsections.

Orbital Taurus Program Manager — The OrbitalProgram Manager has direct responsibility andaccountability for the Taurus Program at Orbital,including all elements required to provide areliable and timely launch service and to ensurecustomer satisfaction. The Program Manager’sresponsibilities include management ofschedules, budgets and deliverables; customerinterface; vehicle engineering; change control;manufacturing; and integration and launchoperations.

Taurus Mission Manager — The missionintegration management structure is illustrated inFigure 7-3. The Taurus Mission Manager is thesingle point of contact for all aspects of a specificmission. This person has overall programauthority and responsibility to ensure that payloadrequirements are met and that the appropriatelaunch services are provided. The Taurus MissionManager will chair the Mission IntegrationWorking Groups (MIWGs). The MissionManager’s responsibilities include detailedmission planning, launch vehicle production

• Lead Mechanical Engineer• Lead Electrical Engineer• Lead Systems Engineer• Lead Ordnance Engineer• Lead GN&C Engineer• Lead Propulsion Engineer

• Mission Manager, Flt. 1• Mission Manager, Flt. 2• Mission Manager, Flt. n

• Safety • Quality• Configuration Management

Taurus Program Manager/Deputy Program Manager

ContractsFlight Assurance Manager

SubcontractsFinance

Figure 7-1. Taurus Program Organization.UG061

Manager, SiteOperations

Manager, Chandler

Operations

ChiefProgramEngineer

Director,Launch

Services



coordination, payload integration services,systems engineering, mission-peculiar design andanalyses coordination, payload interface

definition, launch range coordination, integratedscheduling, launch site processing, and flightoperations.

Taurus Chief Program Engineer/EngineeringLeads — The Orbital engineering activities withinthe Taurus Program are directed by the TaurusChief Program Engineer. The Chief ProgramEngineer is supported by a team of engineeringleads and support staffs representing all of therequired engineering disciplines. This group ofindividuals is responsible for supporting missionintegration activities for all Taurus missions. Theprimary mission engineering support tasks includemission analyses, software development, mission-peculiar hardware design and testing, andmission-peculiar analyses.

Manager, Site Operations — The Manager ofSite Operations is responsible for supportingoperations for all Taurus missions. Primarymission responsibilities include overall integrationplanning, vehicle procedure development, andrange and production facility coordination.

Vehicle Engineer — The Orbital Vehicle Engineeris responsible for the integration of a specificmission. This person has overall technicalprogram authority and responsibility to ensurethat a vehicle is integrated to support a specificlaunch. The Orbital Vehicle Engineer supportsthe Taurus Mission Manager to ensure that vehiclepreparation is on schedule.

All work that is scheduled to be performed on thevehicle is directed and approved by the VehicleEngineer. This includes preparation andexecution of work procedures, launch vehicleprocessing, and control of hazardous operations.All hazardous procedures are approved by thelaunch range safety representative, the OrbitalVehicle Engineer, and the Orbital Safety Managerprior to execution. In addition, Orbital Safetyand Quality Assurance engineers are alwayspresent to monitor critical and hazardousoperations. Scheduling of payload encapsulationand integration with the launch vehicle and allrelated activities are also coordinated by theVehicle Engineer.

Primary Mission Management Responsibility • Payload Integration • Integrated Scheduling • Mission Planning • Range Coordination • Launch OperationsSingle Point of Contact for Specific MissionCoordinates System Engineering Tasks

TaurusMission Management

Launch Site ProcessingVehicle Integration and TestIntegrated Procedures Preparation and Approval

Vehicle Engineer

Launch Site SchedulingHazardous Operations ControlRange InterfaceFacilities Maintenance and Control

Manager,Site Operations

Mission Orbital AnalysisMission Peculiar Hardware Design/TestMission Peculiar AnalysisVehicle Production PlanningVehicle UpgradesMechanical AnalysesElectrical Analyses

Engineering

Overall Taurus Program ManagementOverall Launch Services ManagementContract AdministrationProgram ControlFlight Assurance • Safety • Reliability • Quality Assurance • Vehicle Configuration Management

TaurusProgram Manager

Figure 7-2. Orbital Mission Responsibilities.UG062



7.2.1 Customer Mission Responsibilities

The customer mission responsibilities aresummarized in Figure 7-4. The Payload MissionManager is responsible for ensuring that theinterests of the mission are best served. Once themission services are defined, the Payload MissionManager has insight into the mission integrationanalyses and tests that are performed by Orbital.In addition, the Payload Mission Manager remainsfully apprised of launch vehicle production andtesting activities as well as launch site activitiesprior to payload arrival. Once the payload isdelivered to the launch site, all activities arecontrolled by an integrated processing schedulewhich is developed by the Mission Managers andapproved by all organizations prior toimplementation.

7.3 Working Groups

The core of the mission integration processconsists of a series of Mission Integration andRange Working Groups (MIWG and RWG,respectively). The MIWG has responsibility forall physical interfaces between the payload andthe launch vehicle. As such, the MIWG createsand implements the Payload-to-Taurus ICD in

addition to all mission-unique analyses, hardware,software, and integrated procedures. The RWGis responsible for the areas of launch siteoperations; range interfaces; safety review andapproval; and flight design, trajectory, andguidance. Documentation produced by the RWGincludes all required range and safety submittals.

Customer Interface

Integrated Procedures DevelopmentVehicle Integration and TestPayload ProcessingLaunch Site Processing

TrajectoryLoadsThermalVentingCoupled LoadsEMI/EMCAcoustic/Vibe/Shock

Range Asset SchedulingFacility InterfaceFacility Activation/CoordinationOversight of Technicians andTest Engineers

All Mission-Unique Analysis:

Production PlanningMission Data Load DevelopmentDesign/Development of Mission-Unique Equipment and Software

Mission Schedule/Budget

Interface Verification Plan and Procedures

ICD Development and Maintenance

Mission Manager

Engineering Manager, Site OperationsVehicle Engineer

Figure 7-3. Mission Integration Management Structure.UG063

Primary Mission Management Responsibility

Single Point of Contact for Specific Mission

Interface Between Payload Office and Orbital

Defines Payload Requirements

Defines Mission Specific Range Requirements

Coordinates Integrated Scheduling

Coordinates Mission Integration Analyses and Test • Mission Orbital Analysis • Mission Peculiar Analysis • Mission Peculiar Design and Test

Figure 7-4. Customer Mission Responsibilities.UG078

PayloadMission Manager



Working Group membership consists of theMission Manager and representatives from Taurus’engineering and operations organizations, aswell as their counterparts from the customerorganization. While the number of meetings,both formal and informal, required to developand implement the mission integration processwill vary with the complexity of the spacecraft,bi-monthly meetings are typical.

7.4 Mission Planning and Development

Orbital will assist the customer with missionplanning and development associated with Tauruslaunch vehicle systems. These services includeinterface design and configuration control,development of integration processes, launchand launch vehicle related analyses, facilitiesplanning, launch campaign planning to includerange services and special operations, andintegrated schedules.

The procurement, analysis, integration and testactivities required to place a customer’s payloadinto orbit are typically conducted over a 26month long standard sequence of events calledthe Mission Cycle. This cycle begins 24 monthsbefore launch, and extends to eight weeks afterlaunch. The duration of the Mission Cycle can beadjusted to meet customer constraints.

Once contract authority to proceed is received,the Mission Cycle is initiated. The contractdesignates the payload, launch date, and basicmission parameters. In response, the TaurusProgram Manager designates an Orbital MissionManager who ensures that the launch service issupplied efficiently, reliably, and on-schedule.

The typical Mission Cycle interweaves thefollowing activities:

a. Mission management, document exchanges,meetings, and formal reviews required tocoordinate and manage the launch service.

b. Mission analyses and payload integration,document exchanges, and meetings.

c. Design, review, procurement, testing andintegration of all mission-peculiar hardwareand software.

d. Range interface, safety, and flight operations

activities, document exchanges, meetings andreviews.

Figure 7-5 details the typical Mission Cycle for aspecific launch and how this cycle folds into theOrbital vehicle production schedule with typicalpayload activities and milestones. A typicalMission Cycle is based on a 24 month intervalbetween mission authorization and launch. Thisinterval reflects the most efficient schedule basedon Orbital’s Taurus and Pegasus programexperience. However, Orbital is flexible tonegotiate either accelerated cycles, which takeadvantage of the Taurus/Pegasus multi-customerproduction sets, or extended cycles required byunusual payload requirements, such as extensiveanalysis or complex payload-launch vehicleintegrated designs or tests.

7.5 Reviews

During the integration process, reviews are heldto provide the coordination of mission participantsand management outside of the regular contactof the Working Groups. Due to the variability incomplexity of different payloads and missions,the content and number of these reviews can betailored to customer requirements. At a minimum,Orbital will conduct two readiness reviews asdescribed below.

Mission Readiness Review — Conducted withinone month of launch, the Mission ReadinessReview (MRR) provides a pre-launch assessmentof integrated launch vehicle/payload/facilityreadiness prior to committing significant resourcesto the launch campaign.

Launch Readiness Review — The LaunchReadiness Review (LRR) is conducted at L-1 dayand serves as the final assessment of missionreadiness prior to activation of range resourceson the day of launch.

7.6 Documentation

Integration of the payload requires detailed,complete, and timely preparation and submittalof interface documentation. As the launch serviceprovider, Orbital is the primary communicationpath with support agencies, which include—but



are not limited to—the various Range supportagencies and U.S. Government agencies such asthe FAA, U.S. Department of Transportation andU.S. State Department. Customer-provideddocuments represent the formal communicationof requirements, safety data, system descriptions,and mission operations planning. The major

products and submittal times associated withthese organizations are divided into two areas—those products that are provided by the customer,and those produced by Orbital.

7.7 Customer-Provided Documentation

Documentation produced by the customer is

Figure 7-5. Typical Master Schedule. UG060

Task

ATP

L-24

L-23

L-22

L-21

L-20

L-19

L-18

L-17

L-16

L-15

L-14

L-13

L-12

L-11

L-10

L-9

L-8

L-7

L-6

L-5

L-4

L-3

L-2

L-4

L-3

L-2

L-9

L-8

L-7

L-6

L-5

L-4

L-3

L-2

L-1

L+1

L+2

L+Weeks

WG #1 WG #2 WG #3 WG #4 WG #5WG #6

WG #7Released

PreliminaryThermal

Final

WG #8

WG #9

PRD, MissionAnnex

Draft

PMAFMA

RF Link

Final L+8 Weeks

FinalPreliminary

PSP ODORPI

Planning

L-Months L-Weeks L-Days

Design/Procurement Pre-Integration Launch-Site Operations

Mission Integration MIWG/RWG ICD ICD VerificationMission Analyses PMA FMA Coupled Loads Integrated Thermal RF Link & CompatibilityRange & Safey Documentation Launch Vehicle Range Response Integrated MSPSP Flight Plan/TrajectoryRocket Motor Production Long Lead Procurement Production Buy-Off and ShipmentVehicle Production Avionics Components Harness Manufacturing Interstage Production Fairing Production Fairing Test TPS Application Fairing Final Assembly Interstage Final Assembly Avionics Testing Ship to FieldVehicle MAB Processing Receive Motors Stage Mechanical Build-Up Receive/Install Avionics System Testing Stage MatesVehicle Payload Integration Final Integrated Procedures Taurus Hardware Ships to PPF Payload Mate to PAF Fairing Installation Payload Transport to SiteLaunch Site Operations Transport Stages to Site System Testing Receive ECE from PPF Flight Simulation #3 ECE Mate, Closeouts, Stacking Final Testing Mission Rehearsal LaunchPost Launch Quick Look Final Flight Report



summarized in Figure 7-6 and detailed in thefollowing paragraphs.

7.7.1 Payload Questionnaire

The Payload Questionnaire is designed to providethe initial definition of payload requirements,interface details, launch site facilities, andpreliminary safety data to Orbital. The customershall provide a response to the PayloadQuestionnaire form (Appendix A) as soon as thespacecraft definition is reasonably firm (e.g.,when the RFP is released) but not later than oneweek after authority to proceed has been provided.The customer’s responses to the payloadquestionaire define the most current payloadrequirements and interfaces and are instrumentalin Orbital’s preparation of numerous documentsincluding the ICD, Preliminary Mission Analysis,and launch range documentation. Additionalpertinent information, as well as preliminarypayload drawings, should also be included withthe response. Orbital understands that a definitiveresponse to some questions may not be feasible.These items are defined during the normal missionintegration process.

7.7.2 Payload Finite Element Model

A payload mathematical model is required foruse in Orbital’s preliminary coupled loadsanalyses. Acceptable forms include either aCraig-Bampton model to 120 Hz or a NASTRANmodel limited to 50,000 DOF.

7.7.3 Safety Documentation

For each Taurus mission, Orbital acts as theinterface between the mission and Range Safety.In order to fulfill this role, Orbital requires safetyinformation from the payload. For launches fromeither the Eastern or Western Ranges, EWR 127-1 provides detailed range safety regulations. Toobtain approval to use the launch site facilities,specified data must be prepared and submitted tothe Taurus Program Office. This informationincludes a description of each payload hazardoussystem and evidence of compliance with safetyrequirements for each system. Drawings,schematics, and assembly and handlingprocedures, including proof test data for all lifting

equipment, as well as any other information thatwill aid in assessing the respective systems shouldbe included. Major categories of hazardoussystems are ordnance devices, radioactivematerials, propellants, pressurized systems, toxicmaterials, cyrogenics, and RF radiation.Procedures relating to these systems as well asany procedures relating to lifting operations orbattery operations should be prepared for safetyreview submittal. Orbital will provide thisinformation to the appropriate safety offices forapproval.

7.7.4 PPF Requirements

The customer is required to provide data on allpayload activities to be performed at the PPF.This includes detailed information of all facilities,services, and support requested by Orbital andthe PPF provider.

7.7.5 Program Requirements Document (PRD)Mission Specific Annex Inputs

To obtain range support from ER and WR, a PRDmust be prepared. A Taurus PRD has beensubmitted and approved by the range. Thisdocument describes requirements needed tosupport the Taurus launch vehicle. For eachlaunch, an annex is submitted to specify therange support needed to meet the mission’srequirements. This annex includes all payloadrequirements as well as any additional Taurus

Payload QuestionnairePayload Finite Element ModelPayload Safety DocumentationPPF RequirementsPayload Inputs to PRD Mission AnnexPayload Input for Acoustic AnalysisPayload Thermal ModelTest Verified Finite Element ModelPayload Input for Launch ORPayload Launch Site ProceduresPayload Inputs to Integrated Test ProceduresFinal Payload Mass PropertiesICD Verification Documentation

ATP + 1 WeekATP + 1 MonthL-15 MonthsL-14 MonthsL-13 MonthsL-13 MonthsL-11 MonthsL-9 MonthsL-6 MonthsL-5 MonthsL-4 Months

L-2 MonthsL-2 Weeks

Deliverable Delivery Date

UG

012

Figure 7-6. Customer-Provided Documentation.



requirements that may arise to support a particularmission. The customer completes all appropriatePRD forms for submittal to Orbital.

7.7.6 Payload Vibroacoustic Model

An acoustic model is required for use in Orbital’smission-specific vibroacoustic analysis. Theanalysis is performed using a customer providedstatistical energy model of the payload in VAPEPSformat or, alternatively, Orbital will construct asuitable model using customer supplied geometricdata, cross-sectional properties, and massproperties of the payload.

7.7.7 Payload Thermal Model for IntegratedThermal Analysis

A payload thermal model is required for use inOrbital’s integrated thermal analysis. The analysisis conducted for three mission phases:

(1) Prelaunch ground operations;(2) Ascent from liftoff until fairing jettision; and(3) Fairing jettison through payload deployment.

Models must be provided in SINDA format.There is no limit on model size although turn-around time may be increased for excessivelylarge models.

7.7.8 Test Verified Payload Finite Element Model

A test verified payload model is required for usein Orbital’s final coupled loads analysis.Acceptable forms include either a Craig-Bamptonmodel to 120 Hz or a NASTRAN model limitedto 50,000 DOF.

7.7.9 Launch Operations Requirements (OR)Inputs

To obtain range support for the launch operationand associated rehearsals, an OR must beprepared. The customer must provide all payloadlaunch day requirements for incorporation intothe mission OR.

7.7.10 Payload Launch Site Test Procedures

All hazardous procedures to be conducted at thelaunch facility must be prepared for safetyapproval, as discussed in Section 7.7.3. In

addition, any payload procedures that areperformed after fairing encapsulation must bepresented to Orbital for review prior to use.

7.7.11 Payload Integrated Test Procedure Inputs

For each mission, Orbital requires detailedspacecraft requirements for integrated activities.With these requirements, Orbital will producethe procedures discussed in Section 7.8.14.

7.7.12 Payload Final Mass Properties

Accurate mass properties are required two monthsprior to launch so that Orbital may submit thefinal trajectory analysis to the range no later than45 days prior to launch. Since the “final”spacecraft mass properties may not be determineduntil after propellant loading or other operationsat the launch site, the requirement for the massproperties provided to Orbital is that they bewithin the specified ICD tolerances to those massproperties measured after final launch siteoperations.

7.7.13 Taurus/Payload ICD VerificationDocumentation

Orbital conducts a rigorous verification programto ensure all requirements on both sides of theinterface have been successfully fulfilled. As partof the ICD, Orbital produces a verification matrixthat indicates how each ICD requirement will beverified (e.g., test, analysis, demonstration, etc.).As part of the verification process, Orbital willprovide the customer with a form to complete foreach interface that is the responsibility of thepayload to meet. The form clearly identifies thedocumentation to be provided as proof ofverification. Likewise, Orbital provides similardata for all interfaces that are the responsibility ofthe Taurus vehicle, as discussed in Section 7.8.18.

7.8 Orbital Produced Documentation

Documentation produced by Orbital issummarized in Figure 7-7 and detailed in thefollowing paragraphs.

7.8.1 Payload Separation Analysis

Orbital performs the necessary evaluations using



the Taurus STEP 6-DOF simulation and theDynamic Analysis and Design System (DADS)tools to ensure that the payload is in the desiredorientation for successful separation at the end ofboost. This analysis verifies that sufficientseparation distance exists between the payloadand the final stage and will include the effects ofseparation system operation and residual finalstage thrust. Attitude and rate errors are definedas part of this process. For customers using theOrbital supplied separation system, the expectedspacecraft tip off rates are also provided.

7.8.2 PPF Requirements Document

Orbital will prepare the necessary documentationto secure the use and required services of anappropriate PPF. The documentation is based oninputs received from the customer.

7.8.3 Program Introduction (PI)

The PI provides the initial notification to thelaunch range that a mission is beginning. It

identifies the general requirement and schedulefor the mission. The resulting response from thelaunch range (Statement of Capabilities) describesthe range's ability to meet the missionrequirements. Based on inputs from the payloadpayload questionnaire, Orbital will produce thePI and submit the PI to the range. If agreeable tothe range, the PI can be accomplished by a shortpresentation. It is recommended that payloadrepresentatives be present at such a meeting.

7.8.4 Coupled Loads Analysis (Preliminary andFinal)

Orbital has developed and validated finite elementstructural models of the Taurus vehicle for use incoupled loads analyses with Taurus payloads.Orbital will incorporate the customer-providedpayload model into the Taurus finite elementmodel and perform a preliminary coupled loadsanalysis to determine the maximum responses ofthe entire stack under transient loads. Once a testvalidated spacecraft model has been delivered to

Preliminary Separation Analysis

PPF RequirementsProgram IntroductionPreliminary Coupled Loads AnalysisPreliminary Mission Analysis (PMA)

Released Interface Control Document

RF Compatibility Analysis

Vibroacoustic AnalysisIntegrated MSPSPProgram Requirements Mission AnnexIntegrated Thermal AnalysisFinal Coupled Loads AnalysisFairing Clearance Analysis

Operations RequirementsIntegrated Launch Site Test ProceduresMission Constraints DocumentCountdown ProcedureFinal Mission AnalysisTaurus ICD Verification DocumentationQuick Look ReportFinal Flight Report

Payload QuestionnairePayload DrawingsPayload PPF Requirements InputPayload QuestionnaireFinite Element ModelPayload QuestionnaireDraft ICDPayload QuestionnairePayload DrawingsMIWGPayload QuestionnairePayload DefinitionStatistical Energy Model of PayloadPayload Safety PackagePayload PRD FormsSINDA or Equivalent Payload ModelTest Verified FEMFinal Coupled Loads AnalysisTaurus Structural Test ResultsPayload InputsPayload Procedure InputsPayload Mission ConstraintsPayload InputsFinal Mass PropertiesNo Payload Input RequiredLaunchRange Data

ATP + 1 Month

L-23 MonthsL-23 MonthsL-19 MonthsL-15 Months

L-14 Months

L-12 Months

L-11 MonthsL-11 MonthsL-9 MonthsL-9 MonthsL-7 MonthsL-6 Months

L-4 MonthsL-3 MonthsL-2 MonthsL-1 MonthL-2 WeeksL-2 WeeksL+3 DaysL+8 Weeks

Deliverable Dependent On Delivery Date

UG

013

Figure 7-7. Orbital Produced Documentation.



Orbital, a final load cycle is completed. In bothcases, results are provided to the customer in atimely fashion after each load cycle is complete.In some cases, interim results can be madeavailable. Close coordination between thecustomer and the Taurus program office isessential to ensure correct output format andwork schedule for the analysis.

7.8.5 Preliminary Mission Analysis

Orbital performs preliminary mission analysesusing the industry standard POST and Orbital-developed STEP simulation tools. The primaryobjective is to determine the compatibility of thepayload with Taurus and to provide succinct,detailed mission requirements, such asperformance capability, accuracy estimates andpreliminary mission sequencing. The results ofthese analyses are incorporated in the MissionICD.

7.8.6 Interface Control Document (ICD)

The Taurus-to-Payload ICD documents allmission-peculiar requirements agreed upon byOrbital and the customer. The ICD contains thepayload description, electrical and mechanicalinterfaces, fairing requirements, targeting criteria,description of the mission-peculiar vehicle, anda description of special LSE and facilities required.As part of this document, Orbital provides acomprehensive matrix which lists all ICDrequirements and the method in which they arevalidated. The initial issue of the mission ICD isbased upon data provided in the PayloadQuestionnaire and initial meetings between thetwo organizations. The ICD is updated asrequired. As a subset of the ICD, the PayloadMechanical Interface Control Drawing (MICD)and Electrical Interface Control Drawing (EICD)are described in the following paragraphs.

7.8.6.1 Payload Mechanical Interface ControlDrawing (MICD)

The Payload MICD is generated by Orbital usingthe customer-provided configuration drawings.This drawing combines the payload and pertinentTaurus components and illustrates all mechanicalinterfaces. The Payload MICD shows the payload-fairing dynamic envelope, payload access door

locations, and other interface details. The payloadMICD will be accepted as part of the ICD betweenOrbital and the customer.

7.8.6.2 Payload Electrical Interface ControlDrawing (EICD)

The Payload EICD is generated by Orbital. Likethe MICD, the EICD shows all payload-to-Tauruselectrical interfaces. The EICD includes a wiringdiagram and pin assignments for all power,control, telemetry, and monitor interfacesrequired to support the payload. The PayloadEICD is formally accepted as part of the ICDbetween Orbital and the customer. If the electricalinterface can be adequately illustrated within thesize limitations of the ICD, it will be included aspart of the Taurus-to-Payload ICD and a separateelectrical drawing will not be required.

7.8.7 RF Compatibility Analysis

Orbital performs the analysis necessary to ensureelectromagnetic compatibility of the launchvehicle with the payload and the appropriaterange.

7.8.8 Vibroacoustic Analysis

A mission-specific vibroacoustic analysis usingthe VAPEPS analytic tool is conducted by Orbitalto determine acoustic levels within the fairingand random vibration levels at the payloadinterface using a customer-supplied payloadstatistical energy model. Alternatively, Orbitalwill construct a suitable model using customersupplied geometric data, cross-sectionalproperties, and mass properties of the payload.

7.8.9 Integrated Missile System Pre-LaunchSafety Package (MSPSP)

This information includes an assessment of anyhazards which may arise from mission specificvehicle functions and by the payload, and isprovided as an appendix to the baseline MSPSP.The customer must provide the informationpertaining to the payload. Orbital assesses thecombined vehicle and payload for hazards andprepares a report of the findings. Orbital willthen forward the integrated assessment to theappropriate launch range for approval.



7.8.10 PRD Mission Specific Annex Inputs

Once customer PRD inputs are received, Orbitalreviews the inputs and, upon issue resolution,submits the mission specific PRD annex to therange for approval. The range will respond witha Program Support Plan (PSP) indicating theirability to support the stated requirements.

7.8.11 Integrated Thermal Analysis

Orbital performs thermal analyses using ourPATRAN model to derive form factors to be usedas input into the industry-standard SINDA finitedifference model. Integrated Taurus/payloadthermal analyses are performed using the payloadparameters supplied to Orbital by the customer.

7.8.12 Fairing Clearance Analysis

Combining coupled loads analysis results, Taurusstructural test deflection results, stackuptolerances, and misalignments, Orbital validatesthe adequacy of the payload dynamic envelopewithin the Taurus payload fairing.

7.8.13 OR Inputs

Orbital submits the OR to obtain range supportfor the launch operation and associated rehearsals.Detailed information regarding all aspects oflaunch day, particularly communicationrequirements, are detailed in the OR. Orbitalgenerates the document, solicits comments fromthe customer, and, upon comment resolution,forwards the mission OR to the range for approval.The range generates the Operations Directive(OD) which is used by range support personnelas guidance for providing the launch day service.

7.8.14 Integrated Launch Site Test Procedures

For each mission, Orbital prepares integratedprocedures for various operations that involvethe payload. These include, but are not limitedto: payload mate to the Taurus adapter; fairingencapsulation; encapsulated payload transportto the launch site; flight simulations; and finalvehicle closeouts. Once customer inputs arereceived, Orbital will provide draft proceduresfor review and comment. Once concurrence isreached regarding all comments, final procedures

will be released prior to use. Draft hazardousprocedures must be presented to the appropriatelaunch site safety organization 90 days prior touse and final hazardous procedures are due 30days prior to use.

7.8.15 Mission Constraints Document

This Orbital-produced document summarizeslaunch day operations for the Taurus vehicle aswell as for the payload. Included in this documentis a comprehensive definition of the Taurus andpayload launch operations constraints, theestablished criteria for each constraint, thedecision making chain of command, and asummary of personnel, equipment,communications, and facilities that will supportthe launch.

7.8.16 Final Countdown Procedure

Orbital produces the launch countdownprocedure that readies the Taurus and payloadfor launch. All Taurus and payload activities thatare required to be performed during the finalcountdown are included in the procedure.

7.8.17 Final Mission Analysis

The Final Mission Analysis consists of severaldetailed analyses which thoroughly evaluate theplanned mission and its effects throughoutpowered flight. These analyses include, but arenot limited to: trajectory, injection accuracy, re-contact and CCAM, venting, launch window,guidance, and stability/control analyses. Thisanalysis, along with other analyses and softwareverification efforts, results in a verified mission-unique Mission Data Load used for final flightsimulations performed at the launch site as wellas for launch.

7.8.18 Taurus/Spacecraft ICD VerificationDocumentation

Orbital is responsible for providingdocumentation for all Taurus interfacerequirements. This documentation will beprovided for customer review and approval aspart of the team effort to complete a thoroughverification that ICD requirements have beenmet. For non-U.S. customers, this information is



limited by the rules set forth in Orbital’s TechnicalAssistance Agreement (TAA) from the StateDepartment.

7.8.19 Post-Launch Analyses

Orbital provides post-launch analyses to thepayload in two forms. The first is a quick-lookassessment provided within three days of launch.The quick-look data report includes preliminarytrajectory performance data, orbital accuracyestimates, system performance preliminaryevaluations, and a preliminary assessment ofmission success. The second, a more detailedfinal report of the mission, will be provided to thecustomer within eight weeks of launch. Includedin the final mission report is the actual missiontrajectory, event times, significant events,environments, orbital parameters and otherpertinent data reduced from on-board telemetryand range tracking sensors. Photographic andvideo documentation, as available, is includedas well. Orbital also analyzes telemetry datafrom each launch to validate Taurus’ performanceagainst the mission ICD requirements. In thecase of any mission anomaly, Orbital will conductan investigation and closeout review.

Section 8Optional Services


Taurus User's Guide Section 8.0—Optional Services

8.1 Optional Services

Orbital offers a variety of optional services tosatisfy the varied requirements of Taurus payloads.This section of the User’s Guide details theoptional services that are currently available toTaurus customers. If a unique requirement existsthat is not addressed within this section, pleaseinquire directly to the Taurus Program todetermine how best to meet the need.

8.2 Dual Manifest Options

The Taurus Program possesses significantexperience in implementing dual-manifestconfigurations. Two of the three launchesconducted to date have delivered multiplespacecraft to orbit, as will two of the next threemissions on the manifest. Orbital provides twopayload support structures that enable dualpayload manifesting while maintainingindependent load paths for the payloads. The63” (1.6 m) diameter Dual Payload Attach Fitting(DPAF) was successfully flown on the secondTaurus mission. The 50” (1.3 m) diameter AftPayload Capsule (APC) will fly on the fourth andsixth Taurus missions.

8.2.1 63” Dual Payload Attach Fitting

The 63" DPAF is comprised of a graphite/epoxycylinder, a 63” diameter frangible joint separationsystem, and a graphite/epoxy forward payloadcone. The forward payload cone utilizes thesame basic design as the standard Taurus payloadcone with only minor modifications for use in theDPAF application. The forward payload coneand the cylindrical section carry the loads inducedby the forward payload, while an aft payloadcone supports the loads induced by the aftpayload. A contamination shield is provided toprevent cross-contamination between the twopayload cavities. After the forward payload issuccessfully separated, the forward payload coneis deployed to allow for the successful separationof the aft payload. The deployment sequence isdesigned to ensure collision avoidance amongthe multiple bodies.

The DPAF has been designed and qualified sothat the DPAF length can be tailored to meet

mission-unique requirements. The DPAF cylinderhas been qualified for use in lengths up to 85inches (2.2 m). Figure 8-1 depicts the payloaddynamic envelopes available for the forward andaft payloads as a function of DPAF length. Notethat the diameter of the envelope available forthe secondary payload is a function of payloadseparation clearances, and, as such, is dependenton payload geometry and mass properties andthe separation system that is selected. Typically,the available envelope is approximately 48 inches(1.2 m) in diameter.

8.2.2 50” Aft Payload Capsule

The 50" APC is comprised of two graphite/epoxyshells joined by a 50” diameter frangible joint

Figure 8-1. 92" Fairing with 63" DPAF.UG072

DynamicEnvelope

19.2448

ø

inmm

Dimensions in

X = DPAF Height

Dim X

5.1130

60.01524

130.3-X3310-X

52.51334

ø

34.5876

63" DPAF

ø 80.92055



separation system. The graphite/epoxy shellsisolate the aft payload from the loads induced bythe forward payload. A contamination shield isprovided to prevent cross-contamination betweenthe two payload cavities. After the forwardpayload is successfully separated, the forwardAPC section is deployed to allow for the successfulseparation of the aft payload. The deploymentsequence is designed to ensure collisionavoidance among the multiple bodies.

The APC has been designed and qualified so thatits length can be tailored to meet mission-uniquerequirements. The maximum available length is72 inches (1.8 m). Figure 8-2 depicts the payloaddynamic envelopes available for the forward andaft payloads as a function of APC length. Thediameter of the envelope available for thesecondary payload is a function of payloadseparation clearances, and, as such, is dependenton payload geometry and mass properties andthe separation system that is selected. Typicallythe available envelope is approximately 40 inches(1.0 m) in diameter. Note that the APC may beused with either the 63” or 92” payload fairingconfigurations.

8.3 Payload Fairing Options

Orbital offers a number of non-standard servicesthat can be selected to customize Taurus’ payloadfairings to meet the requirements of a particularmission. These services are detailed in thesubsequent paragraphs.

8.3.1 Additional Fairing Access Doors

Additional access doors can be provided foreither payload fairing configuration. The standarddoor sizes are 12 in x 12 in (305 mm x 305 mm)for the 63” payload fairing and 18 in x 24 in (457mm x 610 mm) for the 92” payload fairing. Aspart of this non-standard service, Orbital willperform structural analyses to verify theacceptability of the mission-specific doorconfiguration.

8.3.2 Fairing Access Doors of Non-StandardSize

Access doors of non-standard size can be providedfor either payload fairing configuration. Includedin this non-standard service are structural analyses

of the payload fairing, design and analysis of theaccess cover and mounting frame, andmodification of the fairing acoustic blanket layout.This non-standard service may also be used toprovide access doors in locations beside thecylindrical section of the payload fairing.

8.3.3 Fairing Acoustic Blanket Tailoring

The thickness of the payload fairing acousticblankets can be tailored to meet mission-uniqueenvironmental requirements. There will be acorresponding impact on vehicle performanceand the available payload fairing dynamicenvelope.

UG073

30762

DynamicEnvelope

X

ø

481219

ø

25635

25635

102.9-X2614-X

541372

ø

48.91242

ø

5.1130

inmm

Dimensions in

X = APC Height

50" APC

Figure 8-2. 63" Fairing with APC.



8.4 Optional Mechanical Interfaces

Orbital offers two alternate separating interfacesand one non-separating interface as optionalservices. These systems are discussed in thefollowing paragraphs along with an optionalsurface treatment for thermal control purposes.

8.4.1 High Capacity Payload Separation System

Orbital can provide a high capacity payloadseparation system with a 38.81 inch (986 mm)diameter payload bolted interface. This systemprovides a significantly higher payload capabilityat the cost of a somewhat lower tip-offperformance and a higher interface shockenvironment. The separation system ismanufactured for Orbital by Saab Ericsson Spaceof Sweden. Saab has extensive experiencesupplying separation systems for a wide range oflaunch vehicles and payloads. The high capacityTaurus 38.81" system is based on a separationsystem with considerable flight heritage.

The separation system is a marmon clamp banddesign that employs two aluminum interfacerings that are clamped by dual, semi-circularstainless steel clamp bands with aluminum clampshoes. Separation velocity is provided by up toeight matched spring actuators that impart up to27.7 ft-lb (37.6 Joules) of energy. The electricalinterface across the separation plane isaccomplished via two low separation impulseconnector assemblies. The structural capacity ofthe separation system is detailed in Figure 8-3.

8.4.2 23” Payload Separation System

Orbital can also provide a payload separationsystem with a 23.25 inch (591 mm) diameterpayload bolted interface as a non-standardservice. Orbital developed this system for thePegasus program to support smaller payloads.

The 23.25" separation system is a marmon clampband design that employs two aluminum interfacerings that are clamped by a titanium clamp bandwith aluminum clamp shoes. Separation velocityis provided by a set of four matched separationsprings that impart 16.4 ft-lbs (22.2 Joules) of

energy. An adapter cone is provided with the23.25" separation system to interface with thestandard Taurus 38.81” interface. The electricalinterface across the separation plane isaccomplished via two connector assemblies.One connector assembly is dedicated for payloadtelemetry and the other for pyro commands.Each assembly includes a MIL-C-38999 Series,Type II connector (with the locking collarremoved) and all required mounting bracketry.The structural capacity of the 23.25" separationsystem is provided in Figure 8-4.

8.4.3 Non-Separating Payload Interface

Orbital can provide a non-separating payloadinterface as a non-standard service. The non-separating interface is a 38.81” diameter boltcircle. In this optional service, the customer willsupply any required separation connectors/harnessing. This harnessing must interface witha connector plate located on the payload supportstructure.

8.4.4 Separation System Thermal Treatment

As a non-standard service, the forward ring of thepayload separation system can be coated withsurface treatments for on-orbit thermal controlpurposes.

Figure 8-3. High Capacity Payload Separation SystemStructural Capability.

UG074



Pay

load

c.g

., in

(m

)

300(7.6)

250(6.4)

200(5.1)

150(3.8)

100(2.5)

50(1.3)

0500

(227)1000(454)

1500(680)

2000(907)

2500(1134)

3000(1360)



8.5 Payload Isolation System

Orbital offers a flight-proven payload isolationsystem as a non-standard service. This mechanicalisolation system has demonstrated the capabilityto significantly alleviate the dynamic loadsinduced by the resonant burn environmentinherent to the Castor 120 solid rocket motor.The isolation system can provide relief to boththe overall payload center of gravity loads andcomponent or subsystem responses. The isolationsystem does impact overall vehicle performance(by approximately 20-40 lb [9-18 kg]) and theavailable payload dynamic envelope (byapproximately 1.0 in [25.4 mm]).

8.6 Performance Enhancements

If more performance is required than can beprovided by the standard Taurus configuration,there are several options for increasing theavailable mass-to-orbit. Using stretched Orionmotors for Stages 1 and 2 (known as "XL" motors)increases Taurus performance to low earth orbitsby approximately 440 lb (200 kg).

In another option, the Orion 38 Stage 3 motorcan be replaced with a spinning STAR 37FMmotor as a non-standard service. The payloaddynamic envelopes provided by this option aredepicted in Figure 8-5 and Figure 8-6 for the 63"

and 92" payload fairings, respectively. Use of theSTAR 37FM spinning upper stage boosts Taurusperformance to low Earth orbits by approximately350 lb (160 kg), in large part because the avionicsstructure is jettisoned once the STAR 37/payloadstack has been spun up. The spinning upperstage option is particularly useful for high energytrajectories, as shown by the performance curvesin Figure 8-7.

Along with the STAR 37FM motor itself, this non-standard service includes a de-spin system, atumble system for collision and contaminationavoidance, a nutation control system, a payloadseparation system, and spin-balancing of thespacecraft and motor assembly.

8.7 Contamination Control Options

Understanding that payloads have varyingrequirements for cleanliness, Orbital offers anumber of contamination control options. Tauruscustomers can select any combination of theseoptions to meet the unique needs of their payloads.

8.7.1 Class 10,000 Integration Environment

As a non-standard service, the payload processingfacility can maintain a Class 10,000 environment.The Class 10,000 environment will be certifiedand maintained for all payload processingoperations up to and including the fairingencapsulation procedure.

8.7.2 Fairing Surface Cleanliness Options

The internal surfaces of the payload fairing andpayload cone structures can be cleaned andcertified to precision cleaning levels. Included inthis non-standard service are development of amission-unique contamination control plan,cleaning of the fairing and payload cone internalsurfaces and components, and certification perMil-Std-1246. Cleanliness levels as stringent asLevel 500A can be provided.

8.7.3 Instrument Purge or Spot Cooling System

Orbital can provide gaseous nitrogen or heliumpurges to the payload. The instrument purgesystem will supply the purge gas to a disconnectfitting on the payload separation system from the

Figure 8-4. 23.25" Payload Separation SystemStructural Capability.

UG075



400(10.2)

300(7.6)

200(5.1)

100(2.5)

0100(45)

200(91)

300(136)

400(181)

500(227)

600(272)

700(317)

800(363)

900(408)

1000(454)

Pay

load

c.g

., in

(m

)



Figure 8-5. Payload Dynamic Envelope for 63" Diameter Fairing, STAR 37FM Configuration.UG-40

ø55.3(ø1,405)

ø48.0(ø1,219)

ø30.3(ø770)

114.5(2,908)

89.5(2,273)

64.5(1,638)

Payload DynamicEnvelope

Fairing Envelope

Payload Envelope

Motor

Payload Separation System

Notes:



3. Orbital Requires Definition of Payload Features Within 2 Inches (50.8 mm) ofPayload Envelope

4. Projections of Payload AppendagesBelow Payload Separation Plane MayBe Permitted But Must Be CoordinatedWith the Taurus Program Office

5. One 12 Inch x 12 Inch(305 mm x 305 mm) Access DoorProvided as Standard Service



ø19.2(488)

154.0(3912)

188.5(4788)

Fairing Envelope

Payload Envelope

Motor

Payload Separation System

Notes:


2. Acoustic Blanket Thickness Is 1 inch (25.4 mm)

3. Orbital Requires Definition of PayloadFeatures Within 2 Inches (50.8 mm ) ofPayload Envelope

4. Projections of Payload AppendagesBelow the Payload Separation PlaneMay Be Permitted but Must Be Coordin-ated with the Taurus Program Office

5. One 18 Inch (457 mm) x 24 Inch(610 mm) Access Door Provided asStandard Service

Figure 8-6. Payload Dynamic Envelope for 92" Diameter Fairing, STAR 37FM Configuration.UG-22

94.0(2388)

Payload DynamicEnvelope

ø52.5(1334)

ø80.9(2055)



time of fairing encapsulation through lift-off. Thegas flow rate and cleanliness can be tailored tomeet mission-unique requirements. This systemcan also be used for cooling a specific locationon the payload.

8.7.4 Hydrocarbon Monitoring

As a non-standard service, Orbital can monitorhydrocarbon levels in the PPF and inside thepayload fairing after encapsulation. The detailedmonitoring requirements will be tailored on amission-unique basis.

8.8 Real-Time Telemetry Downlink ThroughPayload Separation

For some Taurus missions, payload separationoccurs after loss of the telemetry signal from therange. This non-standard service provides real-time telemetry through payload separation usingmobile assets.

8.9 Payload Propellant Loading

Payload loading services for mono-propellanthydrazine or bi-propellant systems can beprovided by Orbital at the PPF.

8.10 On-Board Video System

Orbital can provide an on-board video camerasystem. The camera system provides a real-timevideo downlink from forward- and aft-lookingcameras mounted on the Taurus Stage 2 motor.The cameras switch views (between forward andaft) as commanded by the flight computer tocapture critical staging events and fairingseparation.

UG077a

Figure 8-7. Taurus Performance to High EnergyTrajectories.

Pay

load

Mas

s (k

g)

-2 0 2 4 6 8 10 12 14 16 18 20 22 24 26 28 30

320

300

280

260

240

220

200

180

160

140

120

100


Pay

load

Mas

s (lb

)

-2 0 2 4 6 8 10 12 14 16 18 20 22 24 26 28 30

700

650

600

550

500

450

400

350

300

250

200


C3 (km /s )2 2

C3 (km /s )2 2

Appendix ATaurus Payload Questionnaire

taurus launch system september 1999 - brown …...release 3.0 september 1999 i thank you for...

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