Minotaur User's Guide · This Minotaur User's Guide is intended to familiarize potential space...

Minotaur

User's Guide

Minotaur

User's Guide

Minotaur

User's Guide

March 2002

Release 1.0

Approved for Public Release

Distribution Unlimited

© 2002 by Orbital Sciences Corporation. All rights reserved.

March 2002

Release 1.0

Minotaur® Users Guide

Copyright © 2002 by Orbital Sciences Corporation.All Rights Reserved.

ORBITAL SCIENCES CORPORATION

Approved for Public ReleaseDistribution Unlimited

Minotaur User's Guide Preface

Release 1.0 March 2002

This Minotaur User's Guide is intended to familiarize potential space launch vehicle users with theMinotaur launch system, its capabilities and its associated services. The launch services describedherein are available for US Government sponsored missions via the United States Air Force (USAF)Space and Missile Systems Center, Detachment 12, Rocket Systems Launch Program (RSLP).

Readers desiring further information on Minotaur should contact the USAF OSP Program Office:

USAF SMC Det 12/RP3548 Aberdeen Ave SEKirtland AFB, NM 87117-5778

Telephone: (505) 846-8957Fax: (505) 846-1349

Additional copies of this User's Guide and Technical information may also be requested fromOrbital at:

Business DevelopmentOrbital Sciences CorporationLaunch Systems Group3380 S. Price RoadChandler, AZ 85248

Telephone: (480) 814-6028E-mail: [email protected]


Minotaur User's Guide

1.0 INTRODUCTION ........................................................................................................... 1-1

2.0 MINOTAUR LAUNCH SERVICE...................................................................................... 2-12.1 Minotaur Launch System Overview ......................................................................... 2-12.2 Minotaur Launch Service ......................................................................................... 2-12.3 Minotaur Launch Vehicle ......................................................................................... 2-2

2.3.1 Lower Stack Assembly ................................................................................... 2-22.3.2 Upper Stack Assembly ................................................................................... 2-3

2.4 Launch Support Equipment ...................................................................................... 2-4

3.0 GENERAL PERFORMANCE ............................................................................................. 3-13.1 Mission Profiles ........................................................................................................ 3-13.2 Launch Sites ............................................................................................................. 3-1

3.2.1 Western Launch Sites ................................................................................... 3-13.2.2 Eastern Launch Sites ..................................................................................... 3-2

3.3 Performance Capability ........................................................................................... 3-23.4 Injection Accuracy .................................................................................................. 3-23.5 Payload Deployment .............................................................................................. 3-53.6 Payload Separation ................................................................................................. 3-53.7 Collision/Contamination Avoidance Maneuver ....................................................... 3-6

4.0 PAYLOAD ENVIRONMENT .............................................................................................. 4-14.1 Steady State and Transient Acceleration Loads ......................................................... 4-1

4.1.1 Optional Payload Isolation System................................................................. 4-24.2 Payload Vibration Environment ................................................................................ 4-34.3 Payload Shock Environment ..................................................................................... 4-34.4 Payload Acoustic Environment ................................................................................. 4-34.5 Payload Structural Integrity and Environments Verification ....................................... 4-4

4.5.1 Recommended Payload Testing and Analysis ................................................. 4-44.6 Thermal and Humidity Environments ....................................................................... 4-5

4.6.1 Ground Operations ....................................................................................... 4-54.6.2 Powered Flight .............................................................................................. 4-64.6.3 Nitrogen Purge (non-standard service) ......................................................... 4-7

4.7 Payload Contamination Control ............................................................................... 4-74.8 Payload Electromagnetic Environment ................................................................... 4-7

5.0 PAYLOAD INTERFACES .................................................................................................. 5-15.1 Payload Fairing ........................................................................................................ 5-1

5.1.1 Payload Dynamic Design Envelope ............................................................... 5-15.1.2 Payload Access Door ..................................................................................... 5-15.1.3 Increased Volume Payload Fairing ................................................................. 5-1

5.2 Payload Mechanical Interface and Separation System .............................................. 5-1

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ii

5.2.1 Standard Non-Separating Mechanical Interface ........................................... 5-15.2.2 Separating Mechanical Interface .................................................................. 5-55.2.3 Orbital Supplied Drill Templates .................................................................. 5-5

5.3 Payload Electrical Interfaces ................................................................................... 5-55.3.1 Payload Umbilical Interfaces ........................................................................ 5-55.3.2 Payload Battery Charging ............................................................................. 5-55.3.3 Payload Command and Control .................................................................... 5-55.3.4 Pyrotechnic Initiation Signals ...................................................................... 5-105.3.5 Payload Telemetry ...................................................................................... 5-105.3.6 Non Standard Electrical Interfaces ............................................................. 5-105.3.7 Electrical Launch Support Equipment ......................................................... 5-10

5.4 Payload Design Constraints .................................................................................. 5-105.4.1 Payload Center of Mass Constraints ............................................................ 5-105.4.2 Final Mass Properties Accuracy ................................................................. 5-105.4.3 Pre-Launch Electrical Constraints ............................................................... 5-105.4.4 Payload EMI/EMC Constraints ..................................................................... 5-105.4.5 Payload Dynamic Frequencies ................................................................... 5-115.4.6 Payload Propellent Slosh............................................................................. 5-115.4.7 Payload-Supplied Separation Systems ........................................................ 5-115.4.8 System Safety Constraints ........................................................................... 5-11

6.0 MISSION INTEGRATION ............................................................................................... 6-16.1 Mission Management Approach ............................................................................... 6-1

6.1.1 SMC Det 12/RP Mission Responsibilities ........................................................ 6-16.1.2 Orbital Mission Responsibilities ..................................................................... 6-1

6.2 Mission Planning and Development ......................................................................... 6-26.3 Mission Integration Process ...................................................................................... 6-2

6.3.1 Integration Meetings ...................................................................................... 6-26.3.2 Mission Design Reviews ................................................................................ 6-46.3.3 Readiness Reviews ........................................................................................ 6-4

6.4 Documentation ........................................................................................................ 6-46.4.1 Customer-Provided Documentation ............................................................... 6-4

6.4.1.1 Payload Questionnaire .................................................................... 6-46.4.1.2 Payload Mass Properties .................................................................. 6-46.4.1.3 Payload Finite Element Model ......................................................... 6-56.4.1.4 Payload Thermal Model for Integrated Thermal Analysis .................. 6-56.4.1.5 Payload Drawings............................................................................ 6-56.4.1.6 Program Requirements Document (PRD) Mission Specific

Annex Inputs .................................................................................. 6-56.4.1.6.1 Launch Operations Requirements (OR) Inputs ........................ 6-5

6.5 Safety ....................................................................................................................... 6-56.5.1 System Safety Requirements .......................................................................... 6-56.5.2 System Safety Documentation........................................................................ 6-6

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iii

7.0 GROUND AND LAUNCH OPERATIONS ....................................................................... 7-17.1 Minotaur/Payload Integration Overview ................................................................... 7-17.2 Ground And Launch Operations .............................................................................. 7-1

7.2.1 Launch Vehicle Integration ............................................................................ 7-17.2.1.1 Planning and Documentation ........................................................ 7-17.2.1.2 Vehicle Integration and Test Activities ............................................ 7-1

7.2.1.2.1 Flight Simulation Tests ................................................. 7-17.2.2 Payload Processing/Integration ...................................................................... 7-1

7.2.2.1 Payload to Minotaur Integration .................................................... 7-27.2.2.2 Pre-Mate Interface Testing ............................................................. 7-27.2.2.3 Payload Mating and Verification .................................................... 7-27.2.2.4 Final Processing and Fairing Closeout ........................................... 7-27.2.2.5 Payload Propellant Loading ........................................................... 7-27.2.2.6 Final Vehicle Integration and Test .................................................. 7-2

7.3 Launch Operations .................................................................................................. 7-27.3.1 Launch Control Organization ........................................................................ 7-2

8.0 OPTIONAL ENHANCED CAPABILITIES ......................................................................... 8-18.1 Mechanical Interface and Separation System Enhancements .................................... 8-1

8.1.1 Separation Systems ........................................................................................ 8-18.1.2 Additional Fairing Access Doors .................................................................... 8-18.1.3 Payload Isolation System ............................................................................... 8-18.1.4 Increased Payload Volume............................................................................. 8-1

8.2 Performance Enhancements ..................................................................................... 8-28.2.1 Insertion Accuracy ......................................................................................... 8-2

8.3 Environmental Control Options ................................................................................ 8-28.3.1 Conditioned Air ............................................................................................. 8-28.3.2 Nitrogen Purge .............................................................................................. 8-28.3.3 Enhanced Contamination Control .................................................................. 8-3

8.3.3.1 High Cleanliness Integration Environment (Class 10K or 100K) .......... 8-38.3.3.2 Fairing Surface Cleanliness Options ................................................... 8-38.3.3.3 High Cleanliness Fairing Environment ............................................... 8-3

8.4 Enhanced Telemetry Options ................................................................................... 8-38.4.1 Enhanced Telemetry Bandwidth ..................................................................... 8-38.4.2 Enhanced Telemetry Instrumentation ............................................................. 8-38.4.3 Navigation Data ............................................................................................ 8-4

9.0 SHARED LAUNCH ACCOMMODATIONS ..................................................................... 9-19.1 Load-Bearing Spacecraft .......................................................................................... 9-19.2 Non Load-Bearing Spacecraft ................................................................................... 9-2

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TABLE OF CONTENTS

SECTION PAGE

LIST OF FIGURES

FIGURE PAGE

Figure 2-1 OSP Minotaur Launch Vehicle ......................................................................... 2-1Figure 2-2 OSP Minotaur Launch Vehicle Configuration ................................................... 2-2Figure 2-3 Minotaur Upper Stack Assembly Processing at Orbital's

Vehicle Assembly Building at VAFB ................................................................. 2-3Figure 2-4 Minotaur EGSE Configuration ........................................................................... 2-5Figure 2-5 Minotaur Launch Control Consoles Configuration ............................................ 2-5Figure 3-1 Minotaur Typical Mission Profile ...................................................................... 3-1Figure 3-2 Minotaur Launch Site Options .......................................................................... 3-2Figure 3-3 Minotaur Performance - California Spaceport Launches (SSI CLF) .................... 3-3Figure 3-4 Minotaur Performance - Kodiak Launch Complex Launches ............................ 3-3Figure 3-5 Minotaur Performance - Spaceport Florida Launches ....................................... 3-4Figure 3-6 Minotaur Performance - Virginia Spaceflight Center Launches ......................... 3-4Figure 3-7 Stage Impact Points for Typical Sun-Synchronous Launch From VAFB ............. 3-5Figure 3-8 Injection Accuracies to Low Earth Orbits ........................................................ 3-5Figure 3-9 Typical Pre-Separation Payload Pointing and Spin Rate Accuracies .............. 3-5Figure 4-1 Phasing of Dynamic Loading Events ............................................................... 4-1Figure 4-2 Payload Design CG Net Load Factors (Typical) .............................................. 4-1Figure 4-3 Minotaur 3-Sigma High Maximum Acceleration as a Function

of Payload Weight ............................................................................................ 4-2Figure 4-4 Payload Random Vibration Environment During Flight .................................... 4-3Figure 4-5 Maximum Shock Environment - Launch Vehicle to Payload............................. 4-3Figure 4-6 Maximum Shock Environment - Payload to Launch Vehicle ............................. 4-3Figure 4-7 Payload Acoustic Environment During Liftoff and Flight ................................. 4-4Figure 4-8 Factors of Safety Payload Design and Test ..................................................... 4-4Figure 4-9 Recommended Payload Testing Requirements................................................ 4-5Figure 4-10 Payload Thermal and Humidity Environments ................................................ 4-5Figure 4-11 Minotaur Worst-Case Payload Fairing Inner Surface Temperature

During Ascent (Payload Region) ....................................................................... 4-6Figure 4-12 Minotaur Launch Vehicle RF Emitters and Receivers ........................................ 4-8

APPENDIX A Payload Questionnaire ..................................................................................... A-1

APPENDIX B Electrical Interface Connectors ....................................................................... B-1



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LIST OF FIGURESCONTINUED

FIGURE PAGEFigure 5-1 Payload Fairing Dynamic Envelope with 38 in (97 cm) Diameter

Payload Interface ............................................................................................ 5-2Figure 5-2 Payload Fairing Access Door Placement Zone ................................................. 5-3Figure 5-3 Non-Separable Payload Mechanical Interface .................................................. 5-4Figure 5-4 38 in (97 cm) Separable Payload Interface ...................................................... 5-6Figure 5-5 23 in (59 cm) Separable Payload Interface ...................................................... 5-7Figure 5-6 17 in (43 cm) Separable Payload Interface ...................................................... 5-8Figure 5-7 Payload Separation Velocities Using the Standard

Separation System ............................................................................................ 5-9Figure 5-8 Vehicle/Spacecraft Electrical Connectors and Associated

Electrical Harnesses ......................................................................................... 5-9Figure 5-9 Payload Mass Properties Measurement Tolerance .......................................... 5-10Figure 6-1 OSP Management Structure .............................................................................. 6-1Figure 6-2 Typical Minotaur Mission Integration Schedule ................................................ 6-3Figure 8-1 Softride for Small Satellites (SRSS) Payload Isolation System ............................. 8-1Figure 8-2 Optional 61 in. Diameter Fairing ..................................................................... 8-2Figure 9-1 Typical Load Bearing Spacecraft Configuration .............................................. 9-1Figure 9-2 JAWSAT Multiple Payload Adapter Load Bearing Spacecraft ......................... 9-2Figure 9-3 Five Bay Multiple Payload Adapter Concept .................................................. 9-2Figure 9-4 Dual Payload Attach Fitting Configuration ...................................................... 9-3Figure B-1 Typical Minotaur Payload Electrical Interface Block Diagram ....................... B-2

Release 1.0 March 2002 vi

Minotaur Payload User's Guide Glossary

A Ampere

AADC Alaska Aerospace DevelopmentCorporation

AC Air Conditioning

ACS Attitude Control System

AFRL Air Force Research Laboratory

AODS All-Ordnance Destruct System

ATP Authority to Proceed

C/CAM Collision/Contamination AvoidanceManeuver

CCAS Cape Canaveral Air Station

CDR Critical Design Review

CFE Customer Furnished Equipment

CG,cg Center of Gravity

CLA Coupled Loads Analysis

CLF Commercial Launch Facility

cm Centimeter

CVCM Collected Volatile CondensableMaterials

dB Decibels

deg Degrees

DPAF Dual Payload Attach Fitting

ECS Environmental Control System

EGSE Electrical Ground Support Equipment

EMC Electromagnetic Compatibility

EME Electromagnetic Environment

EMI Electromagnetic Interference

FLSA Florida Spaceport Authority

FM Frequency Modulation

ft Feet

FTLU Flight Termination Logic Unit

FTS Flight Termination System

g Gravitational Force

GACS Ground Air Conditioning System

GFE Government Furnished Equipment

GN2 Gaseous Nitrogen

GPB GPS Position Beacon

GPS Global Positioning System

GSE Ground Support Equipment

GTO Geosynchronous Transfer Orbit

HAPS Hydrazine Auxiliary Propulsion System

HEPA High Efficiency Particulate Air

Hz Hertz

I/F Interface

ICD Interface Control Drawing

ILC Initial Launch Capability

IMU Inertial Measurement Unit

in Inch

INS Inertial Navigation System

ISO International StandardizationOrganization

IVT Interface Verification Test

kbps Kilobits per Second

kg Kilograms

km Kilometer

lb Pound

lbm Pound(s) of Mass

LCR Launch Control Room

LEV Launch Equipment Vault

LITVC Liquid Injection Thrust Vector Control

LOCC Launch Operations Control Center

LRR Launch Readiness Review

LSA Lower Stack Assembly

LSE Launch Support Equipment

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m/s Meters per Second

Mbps Mega Bits per Second

mA Milliamps

MACH Modular Avionics Control Hardware

MDR Mission Design Review

MHz MegaHertz

MIL-STD Military Standard

MIWG Mission Integration Working Group

mm Millimeter

MPA Multiple Payload Adapter

MRD Mission Requirements Document

MRR Mission Readiness Review

MSPSP Missile System Prelaunch SafetyPackage

ms Millisecond

N/A Not Applicable

NCU Nozzle Control Unit

NM Nautical Mile

OD Operations Directive

OD Outside Dimension

ODM Ordnance Driver Module

OR Operations Requirements Document

OSP Orbital Suborbital Program

P/L Payload

PACS Pad Air Conditioning System

PAF Payload Attach Fitting

PCM Pulse Code Modulation

PDR Preliminary Design Review

PI Program Introduction

PID Proportional-Integral-Derivative

POC Point of Contact

ppm Parts Per Million

PRD Program Requirements Document

PSD Power Spectral Density

PSP Prelaunch Safety Package

RCS Roll Control System

RF Radio Frequency

RGIU Rate Gyro Interface Unit

RGU Rate Gyro Unit

rpm Revolutions per Minute

RSLP Rocket Systems Launch Program

RWG Range Working Group

s&a Safe & Arm

scfm Standard Cubic Feet per Minute

SEB Support Equipment Building

sec Second(s)

SINDA Finite Element Thermal Analysis ToolTradename

SLC Space Launch Complex

SLV Space Launch Vehicle

SMC Space and Missile Systems Center

SOC Statement of Capabilities

SPL Sound Pressure Level

SRM Solid Rocket Motor

SRS Shock Response Spectrum

SRSS Softride for Small Satellites

SSI Spaceport Systems International

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STA Station

TLV Target Launch Vehicle

TML Total Mass Loss

TVC Thrust Vector Control

UDS Universal Documentation System

UFS Ultimate Factory of Safety

USAF United States Air Force

V/M Volts per Meter

VAB Vehicle Assembly Building

VAFB Vandenberg Air Force Base

W Watt

WFF Wallops Flight Facility

WP Work Package

YFS Yield Factor of Safety

Release 1.0 March 2002 1-1

Minotaur Payload User's Guide Section 1.0 - Introduction

1.0 INTRODUCTIONThis User’s Guide is intended to

familiarize payload mission planners with thecapabilities of the Orbital Suborbital Program(OSP) Minotaur Space Launch Vehicle (SLV)launch service. This document provides anoverview of the Minotaur system design and adescription of the services provided to ourcustomers. Minotaur offers a variety of enhancedoptions to allow the maximum flexibility insatisfying the objectives of single or multiplepayloads.

The Minotaur’s primary mission is toprovide low cost, high reliability launch servicesto government-sponsored payloads. Minotauraccomplishes this using flight-proven componentswith a significant flight heritage such as surplusMinuteman II boosters, the upper stage Pegasusmotors, the Pegasus Fairing and Attitude ControlSystem, and a mix of Pegasus, Taurus, and sub-orbital Avionics all with a proven, successfultrack record. The philosophy of placing missionsuccess as the highest priority is reflected in thesuccess and accuracy of all Minotaur missions todate.

The Minotaur launch vehicle system iscomposed of a flight vehicle and ground supportequipment. Each element of the Minotaur systemhas been developed to simplify the mission designand payload integration process and to providesafe, reliable space launch services. This User’sGuide describes the basic elements of theMinotaur system as well as optional services thatare available. In addition, this document providesgeneral vehicle performance, defines payloadaccommodations and environments, and outlinesthe Minotaur mission integration process.

The Minotaur system can operate from awide range of launch facilities and geographiclocations. The system is compatible with, andwill typically operate from, commercial spaceportfacilities and existing U.S. Government ranges atVandenberg Air Force Base (VAFB), CapeCanaveral Air Station (CCAS), Wallops FlightFacility (WFF), and Kodiak Island, Alaska. This

User’s Guide describes Minotaur-uniqueintegration and test approaches (including thetypical operational timeline for payloadintegration with the Minotaur vehicle) and theexisting ground support equipment that is used toconduct Minotaur operations.


Minotaur Payload User's Guide Section 2.0 - Minotaur Launch Service

Figure 2-1. OSP Minotaur Launch Vehicle

2. MINOTAUR LAUNCH SERVICE

2.1. Minotaur Launch System OverviewThe Minotaur launch vehicle, shown in

Figure 2-1, was developed by Orbital for theUnited States Air Force (USAF) to provide a costeffective, reliable and flexible means of placingsmall satellites into orbit. An overview of thesystem and available launch services is providedwithin this section, with specific elements coveredin greater detail in the subsequent sections of thisUser’s Guide.

Minotaur has been designed to meet theneeds of United States Government-sponsoredcustomers at a lower cost than commerciallyavailable alternatives by the use of surplusMinuteman boosters. The requirements of thatprogram stressed system reliability,

transportability, and operation from multiplelaunch sites. Minotaur draws on the successfulheritage of three launch vehicles: Orbital’sPegasus and Taurus space launch vehicles andthe Minuteman II system of the USAF. Minotaur’supper two stages and avionics are derived fromthe Pegasus and Taurus systems, providing acombined total of more than 30 successful spacelaunch missions. Orbital’s efforts have enhancedor updated Pegasus and Taurus avionicscomponents to meet the payload-supportrequirements of the OSP program. Combiningthese improved subsystems with the longsuccessful history of the Minuteman II boostershas resulted in a simple, robust, self-containedlaunch system that has been completelysuccessful in its flights to date and is fullyoperational to support government-sponsoredsmall satellite launches.

The Minotaur system also includes acomplete set of transportable Launch SupportEquipment (LSE) designed to allow Minotaur tobe operated as a self-contained satellite deliverysystem. To accomplish this goal, the ElectricalGround Support Equipment (EGSE) has beendeveloped to be portable and adaptable to varyinglevels of infrastructure. While the Minotaursystem is capable of self-contained operationusing portable vans to house the EGSE, it istypically launched from an established rangewhere the EGSE can be housed in available,permanent structures or facilities. This has beenthe case for the first launches from VAFB.

The vehicle and LSE are designed to becapable of launch from any of the four commercialSpaceports (Alaska, California, Florida, andVirginia), as well as from existing U.S. Governmentfacilities at VAFB and CCAS. The Launch ControlRoom (LCR) serves as the actual control center forconducting a Minotaur launch and includesconsoles for Orbital, range safety, and customerpersonnel. Further description of the LaunchSupport Equipment is provided in Section 2.4.

2.2. Minotaur Launch ServiceThe Minotaur Launch Service is provided



Figure 2-2. OSP Minotaur Launch Vehicle Configuration

through the combined efforts of the USAF andOrbital, along with associate contractors includingTRW and Commercial Spaceports. The primarycustomer interface will be with the USAF Spaceand Missile Systems Center, Detachment 12,Rocket Systems Launch Program (RSLP),designated SMC Det 12/RP. Orbital is the launchvehicle provider. For brevity, this integratedteam effort will be referred to as “OSP”. Whereinterfaces are directed toward a particular memberof the team, they will be referred to directly (i.e.“Orbital” or “SMC Det 12/RP”).

OSP provides all of the necessaryhardware, software and services to integrate, testand launch a satellite into its prescribed orbit. Inaddition, OSP will complete all the requiredagreements, licenses and documentation tosuccessfully conduct Minotaur operations. AllMinotaur production and integration processesand procedures have been demonstrated and arein place for future Minotaur missions. The

Minotaur mission integration process completelyidentifies, documents, and verifies all spacecraftand mission requirements. This provides a solidbasis for initiating and streamlining the integrationprocess for future Minotaur customers.

2.3. Minotaur Launch VehicleThe Minotaur vehicle, shown in expanded

view in Figure 2-2, is a four stage, inertiallyguided, all solid propellant ground launchedvehicle. Conservative design margins, state-of-the-art structural systems, a modular avionicsarchitecture, and simplified integration and testcapability, yield a robust, highly reliable launchvehicle design. In addition, Minotaur payloadaccommodations and interfaces have beendesigned to satisfy a wide range of potentialpayload requirements.

2.3.1. Lower Stack AssemblyThe Lower Stack Assembly (LSA) consists

of the refurbished Government Furnished



Figure 2-3. Minotaur Upper Stack AssemblyProcessing at Orbital's Vehicle Assembly

Building at VAFB

Equipment (GFE) Minuteman Stages 1 and 2.Only minor modifications are made to theboosters, including harness interface changes.

The first stage consists of the MinutemanII M55A1 solid propellant motor, Nozzle ControlUnits (NCU), Stage 1 Ignition Safe/Arm, S1/S2Interstage and Stage 1 FTS. Four gimbaled nozzlesprovide three axis control during first stage burn.The Second stage consists of a refurbishedMinuteman II SR19 motor, Liquid Injection ThrustVector Control subsystem (LITVC), S2 ignitionsafe/arm device, a Roll Control System (RCS), andthe Stage 2 FTS components. Attitude controlduring second stage burn is provided by theoperational LITVC and hot gas roll control. ARate Gyro Unit (RGU) is installed on the outerskin of the SR19 to enhance the vehicle controland increase launch availability.

2.3.2. Upper Stack AssemblyThe Minotaur Upper Stack is composed

of Stages 3 and 4 which are the AlliantTechSystems Orion 50XL and 38 SRMs,respectively. These motors were originallydeveloped for Orbital’s Pegasus program andhave been adapted for use on the ground-launched Minotaur vehicle. Common designfeatures, materials and production techniquesare applied to both motors to maximize reliabilityand production efficiency. The motors are fullyflight qualified based on their heritage, designconservatism, ground static fires and over thirtysuccessful flights. Processing of the MinotaurUpper Stack is conducted at the same processingfacility as Pegasus, directly applying theintegration and testing experience of Pegasus tothe Minotaur system (Figure 2-3).

Avionics — The Minotaur avionics systemhas heritage to the Pegasus and Taurus designs.However, the Minotaur design was upgraded toprovide the increased capability and flexibilityrequired by the OSP contract, particularly in thearea of payload accommodations. The flightcomputer, which is common to Pegasus andTaurus, is a 32-bit multiprocessor architecture. Itprovides communication with vehicle

subsystems, the LSE, and the payload, if required,utilizing standard RS-422 serial links and discreteI/O. The Minotaur design incorporates Orbital’sModular Avionics Control Hardware (MACH) toprovide power transfer, data acquisition,Minuteman booster interfaces, and ordnanceinitiation. MACH has exhibited 100% reliabilityon OSP SLV and Target Launch Vehicle (TLV)flights and several of Orbital’s suborbital launchvehicles. In addition, the Minotaur telemetrysystem has been upgraded to provide up to 2Mbps of real-time vehicle data with dedicatedbandwidth and channels reserved for payloaduse.

Attitude Control Systems –– The MinotaurAttitude Control System (ACS) provides three-axis attitude control throughout boosted flightand coast phases. Stages 1 and 2 utilize theMinuteman Thrust Vector Control (TVC) systems.The Stage 1 TVC is a four-nozzle hydraulic system,while the Stage 2 system combines liquidinjection for pitch and yaw control with hot gasroll control. Stages 3 and 4 utilize the same TVCsystems as Pegasus and Taurus. They combinesingle-nozzle electromechanical TVC for pitchand yaw control with a three-axis cold-gasattitude control system resident in the avionicssection providing roll control.

Attitude control is achieved using a three-axis autopilot that employs Proportional-Integral-



Derivative (PID) control. Stages 1 and 2 fly a pre-programmed attitude profile based on trajectorydesign and optimization. Stage 3 uses a set of pre-programmed orbital parameters to place the vehicleon a trajectory toward the intended insertion apse.The extended coast between Stages 3 and 4 is usedto orient the vehicle to the appropriate attitude forStage 4 ignition based upon a set of pre-programmedorbital parameters and the measured performanceof the first three stages. Stage 4 utilizes energymanagement to place the vehicle into the properorbit. After the final boost phase, the three-axis cold-gas attitude control system is used to orient thevehicle for spacecraft separation, contaminationand collision avoidance and downrange downlinkmaneuvers. The autopilot design is modular, soadditional payload requirements such as rate controlor celestial pointing can be accommodated withminimal additional development.

Telemetry Subsystem –– The Minotaurtelemetry subsystem provides real-time healthand status data of the vehicle avionics system, aswell as key information regarding the position,performance and environment of the Minotaurvehicle. This data may be used by Orbital and therange safety personnel to evaluate systemperformance. The minimum data rate is 750 kbps,but the system is capable of data rates up to 2 Mbps.

Payload Fairings –– The baseline Minotaurfairing is identical to the Pegasus fairing design.However, due to differences in vehicle loads andenvironments, the Minotaur implementationallows a larger payload envelope. The Minotaurpayload fairing consists of two composite shellhalves, a nose cap integral to a shell half, and aseparation system. Each shell half is composed ofa cylinder and ogive sections.

Options for payload access doors andenhanced cleanliness are available. A larger 61inch diameter (OD) fairing is also in developmentand is available for future missions. Further detailson the baseline fairing are included in Section 5.1and for the larger fairing in Section 8.1.

With the addition of a structural adapter,either fairing can accommodate multiplepayloads. This feature, described in more detailin Section 9.0 of this User’s Guide, permits two ormore smaller payloads to share the cost of aMinotaur launch, resulting in a lower launch costfor each as compared to other launch options.

2.4. Launch Support EquipmentThe Minotaur LSE is designed to be

readily adaptable to varying launch siteconfigurations with minimal uniqueinfrastructure required. The EGSE consists ofreadily transportable consoles that can behoused in various facility configurationsdepending on the launch site infrastructure. TheEGSE is composed of two primary functionalelements: Launch Control and Vehicle Interface(Figure 2-4). The Launch Control consoles arelocated in a LCR, depending on available launchsite accommodations. The Vehicle InterfaceEGSE is located in structures near the pad,typically called a Support Equipment Building(SEB). Fiber optic connections from the LaunchControl to the Vehicle Interface consoles areused for efficient, high bandwidthcommunications and to minimize the amount ofcabling required. The Vehicle Interface racksprovide the junction from fiber optic cables tothe copper cabling interfacing with the vehicle.

The LCR serves as the control centerduring the launch countdown. The number ofpersonnel that can be accommodated aredependent on the launch site facilities. At aminimum, the LCR will accommodate Orbitalpersonnel controlling the vehicle, two RangeSafety representatives (ground and flight safety),and the Air Force Mission Manager. A typicallayout is shown in Figure 2-5. Mission-unique,customer-supplied payload consoles andequipment can be supported in the LCR andSEB, within the constraints of the launch sitefacilities or temporary structure facilities.Interface to the payload through the Minotaurpayload umbilicals and land lines provides thecapability for direct monitoring of payloadfunctions. Payload personnel accommodationswill be handled on a mission-specific basis.



Figure 2-5. Minotaur Launch Control Consoles Configuration

FTS ControlConsole

Launch Control Room

Serial

Fiber Copper Lines

InterfaceRack

PowerSupplyRack

FTSInterface

Rack

Arming/IgnitionRack

TelemetryMonitor Console

Data ReductionConsole

BackgroundLimit Checking

Console

Flight ComputerConsole

Data DisplayConsole

Power ControlConsole

Data DisplayConsole

Payload I/FRack

PayloadConsole

Support Equipment Building

Patch Panel

TM14025_059

Figure 2-4. Minotaur EGSE Configuration


Minotaur Payload User's Guide Section 3.0 - General Performance

3. GENERAL PERFORMANCE

3.1. Mission ProfilesMinotaur can attain a range of posigrade

and retrograde inclinations through the choice oflaunch sites made available by the readilyadaptable nature of the Minotaur launch system.A typical mission profile to a sun-synchronousorbit is shown in Figure 3-1. High energy andGeosynchronous Transfer Orbit (GTO) missionscan also be achieved. All performance parameterspresented herein are typical for most expectedpayloads. However, performance may varydepending on unique payload or missioncharacteristics. Specific requirements for aparticular mission should be coordinated withOSP. Once a mission is formally initiated, therequirements will be documented in the MissionRequirements Document (MRD). Further detailwill be captured in the Payload-to-Launch VehicleInterface Control Drawing (ICD).

3.2. Launch SitesDepending on the specific mission and

range safety requirements, Minotaur can operatefrom several East and West launch sites, illustratedin Figure 3-2. Specific performance parametersare presented in Section 4.

3.2.1. Western Launch SitesFor missions requiring high inclination

orbits (greater than 60°), launches can beconducted from facilities at VAFB, CA, or KodiakIsland, AK. Both facilities can accommodateinclinations from 60° to 120°, althoughinclinations below 72° from VAFB would requirean out-of-plane dogleg, thereby reducing payloadcapability. Initial Minotaur missions werelaunched from the California Spaceport facility,operated by Spaceport Systems International(SSI), on South VAFB, near SLC-6. The launchfacility at Kodiak Island, operated by the Alaska

Figure 3-1. Minotaur Typical Mission Profile



Aerospace Development Corporation (AADC)has been used for both orbital and suborbitallaunches.

3.2.2. Eastern Launch SitesFor Easterly launch azimuths to achieve

orbital inclinations between 28.5° and 60°,Minotaur can be launched from facilities at CapeCanaveral, FL or Wallops Island, VA. Launchesfrom Florida will nominally use the FloridaSpaceport Authority (FLSA) launch facilities atLC-46 on CCAS, Cape Canaveral, FL. These willbe typically for inclinations from 28.5° to 40°,although inclinations above 35° may havereduced performance due to the need for atrajectory dogleg. The Virginia Spaceflight Centerfacilities at the WFF may be used for inclinationsfrom 30° to 60°. Southeasterly launches fromWFF offer fewer overflight concerns than CCAS.Inclinations below 35° and above 55° are feasible,albeit with doglegs and altitude constraints dueto stage impact considerations.

3.3. Performance CapabilityMinotaur performance curves for circular

and elliptical orbits of various altitudes andinclinations are detailed in Figure 3-3 throughFigure 3-6 for launches from all four Spaceports.These performance curves provide the totalmass above the standard, non-separatinginterface. The mass of any Payload Attach Fitting(PAF) or separation system is to be accounted forin the payload mass allocation. Figure 3-7illustrates the stage vacuum impact points for atypical sun-synchronous trajectory from VAFB.

3.4. Injection AccuracyMinotaur injection accuracy is

summarized in Figure 3-8. Better accuracy canbe provided dependent on specific missioncharacteristics. For example, heavier payloadswill typically have better insertion accuracy, aswill higher orbits. An enhanced option forincreased insertion accuracy is also available(Section 8.2.1). It utilizes the flight-provenHydrazine Auxiliary Propulsion System (HAPS)developed on the Pegasus program.

Figure 3-2. Minotaur Launch Site Options

TM14025_081

WESTERN RANGE

Vandenberg AFB, CA

• Government Launch Sites

• California Spaceport SSI CLF

EASTERN RANGE

Patrick AFB, FL

• Government Launch Sites

• Spaceport Florida

VIRGINIA SPACE FLIGHT

CENTER

Wallops Island, VA

• Commercial Launch Sites

at NASA's Wallops Flight

Facilities

KODIAK LAUNCH

COMPLEX

Kodiak Island, AK



Figure 3-4. Minotaur Performance - Kodiak Launch Complex Launches

TM14025_063a

200 400 600 800 1000 12000

100

200

300

400

500

600

700

800

900

1000

1100

Payload (lbm)

Apo

gee

Alti

tude

(N

M)

Circular Orbit

Elliptical Orbit (Perigee = 100NM)

80 75 70 65

Figure 3-3. Minotaur Performance - California Spaceport Launches (SSI CLF)

200 400 600 800 1000 12000

100

200

300

400

500

600

700

800

900

1000

1100

Payload (lbm)

Apo

gee

Alti

tude

(NM

)

Circular Orbit


99 90 80 72

TM14025_062a



TM14025_065a

300 400 500 600 700 800 900 1000 1100 1200 1300 14000

100

200

300

400

500

600

700

800

900

1000

1100

Payload (lbm)

Apo

gee

Alti

tude

(N

M)

Circular Orbit


50 45 38

Figure 3-6. Minotaur Performance - Virginia Spaceflight Center Launches

Figure 3-5. Minotaur Performance - Spaceport Florida Launches

TM14025_064a

400 500 600 700 800 900 1000 1100 1200 1300 14000

100

200

300

400

500

600

700

800

900

1000

1100

Payload (lbm)

Apo

gee

Alti

tude

(N

M)

Circular Orbit


35 28.5



Figure 3-8. Injection Accuracies to Low EarthOrbits

3.5. Payload DeploymentFollowing orbit insertion, the Minotaur

Stage 4 avionics subsystem can execute a seriesof ACS maneuvers to provide the desired initialpayload attitude prior to separation. Thiscapability may also be used to incrementallyreorient Stage 4 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 specificmission depends upon the spin axis moment ofinertia of the payload and the amount of ACSpropellant needed for other attitude maneuvers.Figure 3-9 provides the typical payload pointingand spin rate accuracies.

3.6. Payload SeparationPayload separation dynamics are highly

dependent on the mass properties of the payloadand the particular separation system utilized.The primary parameters to be considered arepayload tip-off and the overall separation velocity.

TM14025_066

180 170 160 150 140 130 120 110 10020

10

0

10

20

30

40

50

60

70

Longitude (deg)

Latit

ude

(deg

)

Stage 3

Stage 2

Stage 1

Figure 3-7. Stage Impact Points for Typical Sun-Synchronous Launch From VAFB

Figure 3-9. Typical Pre-Separation PayloadPointing and Spin Rate Accuracies

Tolerance Error Type (Worst Case)

Altitude (Insertion Apse) ±10 nm (18.5 km)

Altitude (Non-Insertion ±50 nm (92.6 km) Apse)

Altitude (Mean) ± 30 nm (55.6 km)

Inclination ±0.2°

Error Type Angle Rate

3-Axis Yaw ±1.0° ±0.5°/sec

Pitch ±1.0° ±0.5°/sec

Roll ±1.0° ±0.5°/sec

Spinning Spin Axis ±1.0° ≤10 rpm

Spin Rate 3°/sec



Payload tip-off refers to the angularvelocity imparted to the payload upon separationdue to payload 8 cg offsets and an unevendistribution of torques and forces. If an optionalOrbital-supplied Marmon Clamp-bandseparation system is used, payload tip-off ratesare generally under 4°/sec per axis. Orbitalperforms a mission-specific tip-off analysis foreach payload.

Separation velocities are driven by theneed to prevent recontact between the payloadand the Minotaur upper stage after separation.The value will typically be 2 to 3 ft/sec (0.6 to 0.9m/sec).

3.7. Collision/Contamination AvoidanceManeuver

Following orbit insertion and payloadseparation, the Minotaur Stage 4 will perform aCollision/Contamination Avoidance Maneuver(C/CAM). The C/CAM minimizes both payloadcontamination and the potential for recontactbetween Minotaur hardware and the separatedpayload. OSP will perform a recontact analysisfor post separation events.

A typical C/CAM begins soon afterpayload separation. The launch vehicle performsa 90° yaw maneuver designed to direct anyremaining Stage 4 motor impulse in a directionwhich will increase the separation distancebetween the two bodies. After a delay to allowthe distance between the spacecraft and Stage 4to increase to a safe level, the launch vehiclebegins a “crab-walk” maneuver to impart a smallamount of delta velocity, increasing the separationbetween the payload and the fourth stage of theMinotaur.

Following the completion of the C/CAMmaneuver as described above and any remainingmaneuvers, such as downlinking of delayedtelemetry data, the ACS valves are opened andthe remaining ACS nitrogen propellent isexpelled.


Minotaur Payload User's Guide Section 4.0 - Payload Environment

Supersonic/ Balance of S2 S2 S3 S3 S4Item Liftoff Transonic Max Q S1 Burn Ignition Burn Ignition Burn Burn

Typical FlightDuration

3 sec 17 sec 30 sec 10 sec N/A 70 sec N/A 70 sec 70 sec

Steady State Loads Yes Yes Yes Yes No Yes No Yes Yes

Transient Loads Yes Yes Yes No Yes No Yes No No

Acoustics Yes Yes Yes Yes No No No No No

Random Vibration Yes Yes Yes Yes No Yes No Yes Yes

TM14025_067

4. PAYLOAD ENVIRONMENTThis section provides details of the pre-

dicted environmental conditions that the pay-load will experience during Minotaur groundoperations, powered flight, and launch systemon-orbit operations.

Minotaur ground operations includepayload integration and encapsulation withinthe fairing, subsequent transportation to thelaunch site and final vehicle integration activi-ties. Powered flight begins at Stage 1 ignitionand ends at Stage 4 burnout. Minotaur on-or-bit operations begin after Stage 4 burnout andend following payload separation. To moreaccurately define simultaneous loading andenvironmental conditions, the powered flightportion of the mission is further subdivided intosmaller time segments bounded by criticalflight events such as motor ignition, stage sepa-ration, and transonic crossover.

The environmental design and test crite-ria presented have been derived using measureddata obtained from previous Pegasus, Taurus andMinotaur missions, motor static fire tests, othersystem development tests and analyses. Thesecriteria are applicable to Minotaur configurationsusing the standard 50 in. diameter fairing. Thepredicted levels presented are intended to be rep-resentative of mission specific levels. Missionspecific analyses that are performed as a stan-dard service are documented or referenced inthe mission ICD.

Dynamic loading events that occurthroughout various portions of the flight in-clude steady state acceleration, transient lowfrequency acceleration, acoustic impingement,random vibration, and pyrotechnic shockevents. Figure 4-1 identifies the time phasingof these dynamic loading events and environ-ments and their significance. Pyroshock eventsare not indicated in this figure, as they do notoccur simultaneous with any other significantdynamic loading events.

4.1. Steady State and Transient AccelerationLoads

Design limit load factors due to thecombined effects of steady state and lowfrequency transient accelerations are definedin Figure 4-2. These values include uncertaintymargins and are typical for an 800 lbmpayload.

Figure 4-1. Phasing of Dynamic Loading Events

Figure 4-2. Payload Design CG Net LoadFactors (Typical)

Event Axial (G's) Lateral (G's)

Liftoff 0 ±4.6 0 ±1.6

Transonic 3.2 ±0.5 0.2 ±1.2

Supersonic 3.8 ±0.5 0.4 ±1.1

S2 Ignition 0 ±6.6 0 ±3.3

S3 Ignition 0 ±6.1 0 ±0.5

S3 Burnout See Fig 4-3 0.3 ±0.0

S4 Burnout See Fig 4-3 0.3 ±0.0

TM14025_068



During powered flight, the maximumsteady state accelerations are dependent on thepayload mass. The maximum level can po-tentially occur in either Stage 3 or 4 burn. Fig-ure 4-3 depicts the axial acceleration at burn-out for each stage as a function of payloadmass.

During upper stage burnout, prior tostaging, the transient loads are relatively be-nign. There are significant transient loads thatoccur at both Stage 2 and Stage 3 ignition.During the transient portion of these ignitionevents, the steady state axial loads are rela-tively nonexistent.

As dynamic response is largely governedby payload characteristics, a mission specificCoupled Loads Analysis (CLA) will be per-formed, with customer provided finite elementmodels of the payload, in order to provide moreprecise load predictions. Results will be refer-enced in the mission specific ICD. For pre-

liminary design purposes, Orbital can providepreliminary Center-of-Gravity (CG) netloadsgiven a payload’s mass properties, CG loca-tion and bending frequencies.

4.1.1. Optional Payload Isolation SystemOSP offers a flight-proven payload load

isolation system as a non-standard service.This mechanical isolation system has demon-strated the capability to significantly alleviatethe transient dynamic loads that occur duringflight. The isolation system can provide reliefto both the overall payload center of gravityloads and component or subsystem responses.Typically the system will reduce transient loadsto approximately 50% of the level they wouldbe without the system. In addition, the systemgenerally reduces shock and vibration levelstransmitted between the vehicle and space-craft. The exact results can be expected to varyfor each particular spacecraft and with loca-tion on the spacecraft. The isolation system

TM14025_051

3-S

igm

a H

igh

Max

imum

Axi

al A

ccel

erat

ion

(G's

)

600

1,2001,000800600400200

4.0

5.0

6.0

7.0

8.0

9.0

10.0

11.0

12.0

13.0

14.0

lbm

kg550500450400350300250200150100

S4

S3

50

Payload MassDoes Not Include Random Vibe

Figure 4-3. Minotaur 3-Sigma High Maximum Acceleration as a Function of Payload Weight



Figure 4-4. Payload Random VibrationEnvironment During Flight

Figure 4-5. Maximum Shock Environment -Launch Vehicle to Payload

Figure 4-6. Maximum Shock Enviroment -Payload to Launch Vehicle

PS

D (

g2/H

z)

1e+2

1e+1

1e+3

1e+4

100 1000 10000Frequency (Hz)

TM14025_070

Breakpoints

NaturalFrequency SRS (G)

(Hz) Q=10

100 851000 35001850 5000

10,000 5000

PS

D (

g2/H

z)

1e-2

1e-3

1e-1

10 100 1000 10000Frequency (Hz) TM14025_069

BreakpointsFrequency PSD

(Hz) (g2/Hz)

20 0.00260 0.004

300 0.004800 0.012

1000 0.0122000 0.002

3.5344 gRMS60 Sec Duration

TM14025_052Frequency (Hz)

SR

S (

G's

)

10K

5K

2K

1K

500

200

100

100 200 300 500 1K 2K 3K 5K 10K

50

20

10

Separating ShockNon-Separating Shock

(1000, 3500) (10,000,

3500)

(1850,3000)

(10,000,3000)

(100, 55)

(100, 51)

does impact overall vehicle performance (byapproximately 20-40 lb [9-18 kg]) and theavailable payload dynamic envelope (by up to4.0 in (10.16 cm) axially and approximately 1.0in (2.54 cm) laterally).

4.2. Payload Vibration EnvironmentThe in-flight random vibration curve

shown in Figure 4-4 encompasses all flight vi-bration environments.

4.4. Payload Acoustic EnvironmentThe acoustic levels during lift-off and

powered flight will not exceed the flight limitlevels shown in Figure 4-7. If the vehicle islaunched over a flame duct, the acoustic lev-els can be expected to be lower than shown.This has been demonstrated with flight data.

4.3. Payload Shock EnvironmentThe maximum shock response spectrum

at the base of the payload from all launch ve-hicle events will not exceed the flight limitlevels in Figure 4-5 (separating shock). Formissions that do not utilize an Orbital suppliedpayload separation system, the shock re-sponse spectrum at the base of the payloadfrom vehicle events will not exceed the lev-els in Figure 4-5 (non-separating shock).

If the payload employs a non-Orbitalseparation system, then the shock deliveredto the Stage 4 vehicle interface must not ex-ceed the limit level characterized in Figure4-6. Shock above this level could requirerequalification of components or an accep-tance of risk by the payload customer.



Figure 4-7. Payload Acoustic EnvironmentDuring Liftoff and Flight

10105

110

115

120

125

130

135

100 1000 10000Frequency (Hz)

TM14025_071

Breakpoints

NaturalFrequency SPL (dB-

(Hz) re:2.9e-9psi)

Sou

nd P

ress

ure

Leve

l (dB

- r

e:2.

9e-9

psi)

20253240506380

100125160200250315400

113118123

123.8124.7125.5126.3127.2128

128.8129.7130.5129.9129.3

Breakpoints Cont'd

NaturalFrequency SPL (dB-

(Hz) re:2.9e-9psi)

500630800

1000125016002000250031504000500063008000

10,000

128.6128

126.3124.7123

121.3119.7118

116.3114.7113

111.3109.7108

TM14025_047

Design andTest Options

TestLevel

Dedicated Test Article

Proto-Flight Article

Proof Test EachFlight Article

No Static Test

1.00

1.25

1.10

1.60

1.25

1.50

1.25

2.00

UFS

1.25

1.10

N/A

Design Factor of Safetyon Limit Loads

Yield(YFS)

Ultimate(UFS)

Figure 4-8. Factors of Safety Payload Designand Test

4.5. Payload Structural Integrity and Envi-ronments Verification

The primary support structure for thespacecraft must possess sufficient strength, ri-gidity, and other characteristics required tosurvive the critical loading conditions that ex-ist within the envelope of handling and mis-sion requirements, including worst case pre-dicted ground, flight, and orbital loads. It mustsurvive those conditions in a manner that as-sures safety and that does not reduce the mis-sion success probability. Spacecraft designloads are defined as follows:

a. Design Limit Load — The maximum pre-dicted ground-based, powered flight oron-orbit load, including all uncertainties.

b. Design Yield Load — The Design LimitLoad multiplied by the recommendedYield Factor of Safety (YFS) indicated inFigure 4-8. The payload structure musthave sufficient strength to withstand simul-taneously the yield loads, applied tem-perature, and other accompanying envi-ronmental phenomena for each designcondition without experiencing detrimen-tal yielding or permanent deformation.

c. Design Ultimate Load — The Design LimitLoad multiplied by the recommended Ul-timate Factor of Safety (UFS) indicated inFigure 4-8. The payload structure musthave sufficient strength to withstand simul-taneously the ultimate loads, applied tem-perature, and other accompanying envi-ronmental phenomena without experienc-ing any fracture or other failure mode ofthe structure.

4.5.1. Recommended Payload Testing andAnalysis

Sufficient payload testing and/or analysismust be performed to ensure the safety ofground crews and to ensure mission success.The payload design should comply with thetesting and design factors of safety in Figure4-8. At a minimum, it is recommended thatthe following tests be performed:



Pre-Payload Fairing Installation• Outside VAB Clean Tent• Inside VAB Clean Tent

(Non Standard)

Post-Payload Fairing Installation(GSE)• VAB (GACS)• Transportation to Launch Pad

(Optional)• Lifting Operations (Optional)

PACS (Ground)

23 ±523 ±5

PLF Inlet

23 ±5Ambient(Note 4)Ambient

74 ±1074 ±10

PLF Inlet

74 ±10Ambient(Note 4)Ambient

PLF Inlet13 - 29

PLF Inlet55 - 85

Filtered ACFiltered AC

Filtered ACFiltered Ambient

Filtered AC

ACFiltered AC

(Note 2)

45 ±15<60

(Note 1)Ambient

45 ±1545 ±15

100 K (M6.5)

100 K (M6.5)100 K (M6.5)

Optional

None 100 K (M6.5)

Temp RangeDeg C Deg F

Event Control Humidity (%) Purity Class(Note 3)

Notes:

1. GSE AC Performance is Dependent Upon Ambient Conditions. Temperature Is Selectable and Controlled to Within ±4 ˚F(±2 ˚C) of Set Point. Resultant Relative Humidity is Maintained to 45 ±15%.

2. PACS Performance is Dependent Upon Ambient Conditions (Dew Point). Temperature is Selectable and Controlled Within±4 ˚F (±2 ˚C) of Set Point. Resultant Relative Humidity is Maintained to 45 ±15%.

3. Class 10K (M5.5) Can Be Provided Inside the VAB Clean Tent and at the Payload Fairing Air-Conditioning Inlet on a Mission Specific Basis As a Non-Standard Service.

4. Temperature Control During Transport and Lifting Ops Available as a Non-Standard Service.

Figure 4-9. Recommended Payload TestingRequirements

Figure 4-10. Payload Thermal and Humidity Environments

a. Structural Integrity — Static loads, sinevibration, or other tests should be per-formed that combine to encompass theacceleration load environment pre-sented in Section 4.1.

b. Random Vibration — The flight levelenvironment is defined in Figure 4-4.Recommended test levels are definedin Figure 4-4.

c. Acoustics — Depending on the pay-load configuration, the payload or-ganization may elect to performacoustic testing on the payload, orsub-sections of the payload, in addi-tion to, or in-lieu of, random vibra-tion testing. The acoustic levels aredefined in Figure 4-7.

The payload organization must provideOrbital with a list of the tests and test levels towhich the payload was subjected prior to pay-load arrival at the integration facility.

4.6. Thermal and Humidity EnvironmentsThe thermal and humidity environment

to which the payload may be exposed duringvehicle processing and pad operations are de-fined in the sections that follow and listed inFigure 4-10.

4.6.1. Ground OperationsThe payload environment will be main-

tained by the Ground or Pad Air ConditioningSystems (GACS or PACS). The GACS providesconditioned air to the payload in the VehicleAssembly Building (VAB) after fairing integra-tion. The PACS is used at the launch pad aftervehicle stacking operations. Air Conditioning(AC) is not provided during transport or liftingoperations without the enhanced option thatincludes High Efficiency Particulate Air (HEPA)filtration. The conditioned air enters the fairingat a location forward of the payload, exits aft ofthe payload and is provided up to 5 minutes prior

TestPurposeTest Type Test Level

Random Vibration: the Flight Limit Level Is Characterized inFigure 4-4

Qualification

Acceptance

Protoflight

Flight Limit Level + 6dB

Flight Limit Level

Flight Limit Level + 3dB



TM14025_053a

Flight Time (Sec)

Tem

pera

ture

(˚C

)

Tem

pera

ture

(˚F

)

180

140

100

0 25 50 75 100 125 150

60

20

-20 0

50

100

150

200

250

300

350

• Data Analytically Derived (Using Flight-Verified Models)• Worst Case Heating Profile (Hot Trajectory)• Fairing Inner Surface Temperature in Payload Region

Figure 4-11. Minotaur Worst-Case Payload Fairing Inner Surface Temperatures During Ascent (Payload Region)

to launch. Baffles are provided at the air condi-tioning inlet to reduce impingement velocitieson the payload if required.

Fairing inlet conditions are selected by thecustomer, and are bounded as follows:

• Dry Bulb Temperature: 55-85 °F(13-29 °C) controllable to ±4 °F (±2 °C)of setpoint;

• Dew Point Temperature: 38-62 °F(3-17 °C)

• Relative Humidity: determined by drybulband dewpoint temperature selections andgenerally controlled to within ±15%.Relative humidity is bound by the psy-chrometric chart and will be controlledsuch that the dew point within the fairingis never reached.

4.6.2. Powered FlightThe maximum fairing inside wall tem-

perature will be maintained at less than 200 °F(93 °C), with an emissivity of 0.92 in the region

of the payload. As a non-standard service, a lowemissivity coating can be applied to reduceemissivity to less than 0.1. Interior surfaces aftof the payload interface will be maintained atless than 250 °F (121 °C). Figure 4-11 showsthe worst case transient temperature profile ofthe inner fairing surface adjacent to the pay-load during powered flight.

This temperature limit envelopes themaximum temperature of any component insidethe payload fairing with a view factor to thepayload with the exception of the Stage 4 mo-tor. The maximum Stage 4 motor surface tem-perature exposed to the payload will not ex-ceed 350 °F (177 °C), assuming no shielding be-tween the aft end of the payload and the for-ward dome of the motor assembly. Whether thistemperature is attained prior to payload separa-tion is dependent upon mission timeline.

The fairing peak vent rate is typically lessthan 0.6 psi/sec. Fairing deployment will be



initiated at a time in flight that the maximumdynamic pressure is less than 0.01 psf.

4.6.3. Nitrogen Purge (non-standard service)If required for spot cooling of a payload

component, Orbital will provide GN2 flow tolocalized regions in the fairing as a non-standardservice. The GN2 will meet Grade B specifica-tions, as defined in MIL-P-27401C and can beregulated to at least 5 scfm. The GN2 is on/offcontrollable in the launch equipment vault andin the launch control room.

The system’s regulators are set to a desiredflow rate during prelaunch processing. The sys-tem cannot be adjusted after the launch pad hasbeen cleared of personnel.

Payload purge requirements must be co-ordinated with Orbital via the ICD to ensure thatthe requirements can be achieved.

4.7. Payload Contamination ControlThe Minotaur vehicle, all payload integra-

tion procedures, and Orbital’s contaminationcontrol program have been designed to minimizethe payload’s exposure to contamination fromthe time the payload arrives at the payload pro-cessing facility through orbit insertion and sepa-ration. The Vehicle Assembly Building is main-tained as a visibly clean, temperature and hu-midity controlled work area at all times. TheMinotaur assemblies that affect cleanliness withinthe encapsulated payload volume include thefairing, and the payload cone assembly. Theseassemblies are cleaned such that there is no par-ticulate or non-particulate matter visible to thenormal unaided eye when inspected from 2 to 4feet under 50 ft-candle incident light (VisiblyClean Level II). If required, the payload can beprovided with enhanced contamination controlas a non-standard service.

With enhanced contamination control, asoft walled clean room can be provided to en-sure a FED-STD-209 Class M6.5 (100,000) orClass M5.5 (10,000) environment during all pay-

load processing activities up to fairing encap-sulation. The soft walled clean room andanteroom(s) utilize HEPA filter units to filter theair and hydrocarbon content is maintained at 15ppm or less. The payload organization is respon-sible for providing the necessary clean room gar-ments for payload staff as well as vehicle staffthat need to work inside the clean room.

The inner surface of the entire surface ofthe fairing and payload cone assemblies canbe cleaned to cleanliness criteria which ensuresno particulate matter visible with normal visionwhen inspected from 6 to 18 inches under 100ft-candle incident light. The same will be truewhen the surface is illuminated using black light,3200 to 3800 Angstroms (Visibly Clean Plus Ul-traviolet). In addition, Orbital can ensure thatall materials used within the encapsulated vol-ume have outgassing characteristics of less than1.0% TML and less than 0.1% CVCM. Itemsthat don’t meet these levels can be masked toensure they are encapsulated and will have nosignificant effect on the payload.

With the enhanced contamination controloption, Orbital provides an Environmental Con-trol System (ECS) from payload encapsulationthrough vehicle lift-off. The ECS continuouslypurges the fairing volume with clean filtered air.Orbital’s ECS incorporates a HEPA filter unit toprovide FED-STD-209 Class M5.5 (10,000) air.Orbital monitors the supply air for particulatematter via a probe installed upstream of the fair-ing inlet duct prior to connecting the air sourceto the payload fairing.

Minotaur contamination control is basedon industry standard contamination referencedocuments, including the following:MIL-STD-1246C, “Product Cleanliness Levelsand Contamination Control Program”FED-STD-209E, “Airborne Particulate CleanlinessClasses in Cleanrooms and Clean Zones.”

4.8. Payload Electromagnetic EnvironmentThe payload Electromagnetic Environment



Function Command Tracking Tracking Launch InstrumentationDestruct Transponder Transponder Vehicle on Telemetry

(Optional)

Receive/Xmit Receive Transmit Receive Transmit Transmit

Band UHF C-Band C-Band S-Band S-Band

Frequency 416.5 or 5765 5690 2288.5 2269.5(MHz) 425.0

Bandwidth N/A N/A

Power Output N/A 400 W (peak) N/A 10 W 10 W

Sensitivity -107 dB -70 dB

Modulation Tone Pulse Code Pulse Code PCM/FM PCM/FM

SOURCE 1 2 3 4 5

TM14025_072

Figure 4-12. Minotaur Launch Vehicle RF Emitters and Receivers

(EME) results from two categories of emitters: 1)Minotaur onboard antennas and 2) Range ra-dar. All power, control and signal lines insidethe payload fairing are shielded and properlyterminated to minimize the potential for Elec-tromagnetic Interference (EMI). The Minotaurpayload fairing is Radio Frequency (RF)opaque, which shields the payload from ex-ternal RF signals while the payload is encap-sulated. Based on analysis and supported bytest, the fairing provides 20 db attenuation be-tween 1 and 10000 MHz.

Figure 4-12 lists the frequencies andmaximum radiated signal levels from vehicleantennas that are located near the payloadduring ground operations and powered flight.Antennas located inside the fairing are inac-tive until after fairing deployment. The spe-

cific EME experienced by the payload duringground processing at the VAB and the launchsite will depend somewhat on the specificfacilities that are utilized as well as opera-tional details. However, typically the fieldstrengths experienced by the payload duringground processing with the fairing in placeare controlled procedurally and will be lessthan 2 V/m from continuous sources and lessthan 10 V/m from pulse sources. The highestEME during powered flight is created by theC-Band transponder transmission which re-sults in peak levels at the payload interfaceplane of 28 V/m at 5765 MHz. Range trans-mitters are typically controlled to provide afield strength of 10 V/m or less. This EMEshould be compared to the payload’s RF sus-ceptibility levels (MIL-STD-461, RS03) to de-fine margin.


Minotaur Payload User's Guide Section 5.0 - Payload Interfaces

5. PAYLOAD INTERFACESThis section describes the available me-

chanical, electrical and Launch Support Equip-ment (LSE) interfaces between the Minotaurlaunch vehicle and the payload.

5.1. Payload FairingThe standard payload fairing consists of

two graphite composite halves, with a nosecapbonded to one of the halves, and a separationsystem. Each composite half is composed of acylinder and an ogive section. The two halvesare held together by two titanium straps, both ofwhich wrap around the cylinder section, one nearits midpoint and one just aft of the ogive section.Additionally, an internal retention bolt securesthe two fairing halves together at the surfacewhere the nosecap overlaps the top surface ofthe other fairing half. The base of the fairing isseparated using a frangible joint. Severing thefrangible joint allows each half of the fairing tothen rotate on hinges mounted on the Stage 3side of the interface. A contained hot gas gen-eration system is used to drive pistons that forcethe fairing halves open. All fairing deploymentsystems are non-contaminating.

5.1.1. Payload Dynamic Design EnvelopeThe fairing drawing in Figures 5-1 show

the maximum dynamic envelopes available forthe payload during powered flight. The dynamicenvelopes shown account for fairing and vehiclestructural deflections only. The payload contrac-tor must take into account deflections due tospacecraft design and manufacturing tolerancestack-up within the dynamic envelope. Proposedpayload dynamic envelope violations must beapproved by OSP via the ICD.

No part of the payload may extend aft ofthe payload interface plane without specific OSPapproval. These areas are considered stayoutzones for the payload and are shown in Figure5-1. Incursions to these zones may be approvedon a case-by-case basis after additional verifica-tion that the incursions do not cause any detri-mental effects. Vertices for payload deflectionmust be given with the Finite Element Model to

evaluate payload dynamic deflection with theCoupled Loads Analysis (CLA). The payload con-tractor should assume that the interface plane isrigid; Orbital has accounted for deflections of theinterface plane. The CLA will provide final veri-fication that the payload does not violate the dy-namic envelope.

5.1.2. Payload Access DoorOrbital provides one 8.5 in. x 13.0 in.

(21.6 cm x 33.0 cm), graphite, RF-opaque pay-load fairing access door. The door can be posi-tioned according to user requirements within thezone defined in Figure 5-2. The position of thepayload fairing access door must be defined nolater than L-8 months. Additional access doorscan be provided as a non-standard service.

5.1.3. Increased Volume Payload FairingAn increased volume payload fairing is

currently in development for the Minotaur ve-hicle. This larger fairing is discussed in moredetail in Section 8.1.4.

5.2. Payload Mechanical Interface and Sepa-ration System

Minotaur provides for a standard non-separating payload interface and several optionalOrbital-provided payload separation systems.Orbital will provide all flight hardware and inte-gration services necessary to attach non-separat-ing and separating payloads to Minotaur. Groundhandling equipment is typically the responsibil-ity of the payload contractor. All attachmenthardware, whether Orbital or customer provided,must contain locking features consisting of lock-ing nuts, inserts or fasteners.

5.2.1. Standard Non-Separating Mechanical In-terface

Figure 5-3 illustrates the standard, non-separating payload mechanical interface. Thisis for payloads that provide their own separationsystem and payloads that will not separate. Di-rect attachment of the payload is made on theAvionics Structure with sixty #10 fasteners asshown in Figure 5-3.



Figure 5-1. Payload Fairing Dynamic Envelope With 38 in (97 cm) Diameter Payload Interface

212.983.8

110.043.3

10.04.0

Stayout ZoneClamp/SeparationSystem Components

Payload Interface Connector

φ 38 Inches (97 cm) Payload Separation System

Stayout Zone

HarnessPigtails

to Payload

0°

270°90°

180°

Payload Interface Planefor Non-Separating Payloads

Payload Interface Planefor Payload Separation

System

38 in (97 cm) Avionics Structure22 in (56 cm) Long

119.447.0

ACS StayoutZone

Pyrotechnic Event Connector

77.730.6

+X

+Z

Side View

Forward ViewLooking Aft

PayloadStayout Zones

Legend:

R 270.5 106.5

PayloadDynamicEnvelope

Fairing

100.3 39.5

φ

φ

φ

78.731.0

2.51.0

Dimensions in cmin

Ogive MateLine

TM14025_036

Notes:

(1) Fairing Door Location Is Flexible Within a Specific Region. (Figure 5-2).

(2) Payload Must Request Any Envelope Aft of PayloadInterface Plane.

(3) If Payload Falls within ACSControllability Dead Band They Must Honor ACS Stayout Zone.

(4) If the Payload RequiresNitrogen Cooling, then thePayload Envelope Will beLocally Reduced by 1 InchAlong the Cooling TubeRouting.



Station X+1,749.3 +688.7

AccessDoor Zone

Station X+1,669.8 +657.4

13 5

135

Arc Length

ArcLength

Applies on Either Side of Fairing Joints at0° and 180°

Dimensions in cmin

+X

+Z

Minotaur Coordinates

Notes: 38" Payload

Interface PlaceMinotaur Station X*

(cm/in)

23" PayloadInterface Place

Minotaur Station X*(cm/in)

(1) Entire Access Hole Must Be Within Specified Range.

(2) One 8.5 in x 13.0 in (21.6 cm x 33.0 cm) Door per Mission Is Standard.

(3) Edge of Door Cannot Be Within 5 in (13 cm) of Fairing Joints.

*Without a Soft Ride Load Isolation System

SeparableNon-Separable

1661.4/654.11651.5/650.2

1685.5/663.61677.9/660.6

TM14025_037

Figure 5-2. Payload Fairing Access Door Placement Zone



Payload Harness

MinotaurStage 4Harness

Forward

+X

Harness Access Hole

Forward Interface of φ 38 in (97 cm), 56 cm (22 in) Long Avionics Structure

Rotated 90° CCW Applies at 45° (Pyrotechnic Event) and225° (Payload Interface)

22.99.0

45°

98.638.8Bolt Circle

270°

225°

180°

90°

A

A

119.4 47.0 Fairing Dynamic Envelope

Payload Interface Connector

PayloadStayout

Zone

Pyrotechnic EventConnector

Bolt Circle Consists of0.20 in (60 0.51 cm)Holes Equally Spaced,Starting at 0°

45°

0°

+Y

+Z

10.34.1

MinotaurCoordinates

Forward ViewLooking Aft

φ

φ

Dimensions in cmin

5.7 ±0.092.3 ±0.04

φ

1.90

21.252X

19.952X

TM14025_038

VIEW A-A

Figure 5-3. Non-Separable Payload Mechanical Interface



5.2.2. Separating Mechanical InterfaceThree flight qualified optional separation

systems are available, depending on payloadinterface and size. The 38 in (97 cm) separablepayload interface is shown in Figure 5-4; the23 in (59 cm) separable payload interface isshown in Figure 5-5; the 17 in (43 cm) sepa-rable payload interface is shown in Figure 5-6.Each of these three systems are based on aMarmon band design.

The separation ring to which the payloadattaches is supplied with through holes and theseparation system is mated to the spacecraft dur-ing processing at the VAB. The weight of hard-ware separated with the payload is approxi-mately 8.7 lbm (4.0 kg) for the 38 in (97 cm)system, 6.0 lbm (2.7 kg) for the 23 in (59 cm)system, and 4.7 lbm (2.1 kg) for the 17 in (43cm) system. Orbital-provided attachment boltsto this interface can be inserted from either thelaunch vehicle or the payload side of the inter-face (NAS630xU, dash number based on pay-load flange thickness). The weight of the bolts,nuts, and washers connecting the separationsystem to the payload is allocated to the sepa-ration system and included in the launch ve-hicle mass.

At the time of separation, the payload isejected by matched push-off springs with suffi-cient energy to produce the relative separationvelocities shown in Figure 5-7. If non-standardseparation velocities are needed, differentsprings may be substituted on a mission-specificbasis as a non-standard service.

5.2.3. Orbital Supplied Drill TemplatesOrbital will provide a matched drill tem-

plate to the payload contractor to allow accu-rate machining of the fastener holes. The Or-bital provided drill template is the only approvedfixture for drilling the payload interface. Thepayload contractor will need to send a contractsletter requesting use once they have determinedtheir need dates.

5.3. Payload Electrical InterfacesThe existing design for the payload elec-

trical interface supports battery charging, ex-ternal power, discrete commands, discrete te-lemetry, analog telemetry, serial communica-tion, payload separation indications, and up tofour redundant ordnance events. If an optionalOrbital-provided separation system is utilized,Orbital will provide all the wiring through theseparable interface plane, as illustrated in Fig-ure 5-8 and Figure B-1 of Appendix B. If theoption is not exercised the customer will be re-sponsible to provide the wiring from the space-craft to the separation plane.

5.3.1. Payload Umbilical InterfacesOrbital can provide a maximum of 60

wires from the ground to the spacecraft via adedicated payload umbilical within the vehicle.This internal umbilical is approximately 25 ft inlength. This umbilical is a dedicated pass throughharness, which allows the payload command,control, monitor, and power to be easily con-figured for user requirements. The closest prox-imity for locating customer supplied payloadGSE equipment is the LEV (Launch EquipmentVault). The cabling from the LEV to the launchvehicle is approximately 350 ft.

5.3.2. Payload Battery ChargingOrbital provides the capability for remote

controlled charging of payload batteries, usinga customer provided battery charger. Thispower is routed through the payload umbilicalcable. Up to 4.0 Amps per wire pair can beaccommodated. The payload battery chargershould be sized to withstand the line loss fromthe LEV to the spacecraft.

5.3.3. Payload Command and ControlDiscrete sequencing commands gener-

ated by the launch vehicle’s flight computer areavailable to the payload as closed circuit opto-isolator pulses in lengths of 40 ms multiples. Thecurrent at the payload interface must be lessthan 10 mA. The payload must supply the volt-age source and current limiting resistance tocomplete this circuit.



Dimensions in cmin

Forward View Looking Aft

PayloadPush-OffSprings (4 Places)

Bolt Cutters (2)(Redundant)

Payload Interface

270˚

180˚

0˚

+Z

+Y

MinotaurCoordinates

Payload Pyro Connector

Payload Umbilical Connector

Bolt Circle Consists of0.19 in (60 0.48 cm) Holes Equally Spaced, Starting at 0˚

98.5838.81

10.04.0

5.02.0

Avionics Structure

Payload SeparationClamp Band

Payload InterfacePlaneSeparation

Plane

+Y

+X

4.0 kg (8.7 lbm) Remains with Payload (Includes Harness)

Side View

φ Bolt Circle

TM14025_039

90˚

Figure 5-4. 38 in (97 cm) Separable Payload Interface



180°

0°Payload Push-OffSprings (4 Places)

Payload Interface

Clamp Band


Adpater Cone

Payload UmbilicalConnector

90°


Payload Pyro Connector

Bolt Circle Consists of0.25 in (32 0.64 cm) Holes Equally Spaced, Starting at 0°

+Z

+Y

MinotaurCoordinates

270°

Side View

Adapter Cone


Payload Separation Clamp Band

Payload Attachment

Plane

Bolt Circle

3.751.48

Separation

Plane

2.7 Kg (6.00 lbm)

Remains with Payload

(Includes Harness)

7.492.95

59.0623.25

φ

Dimensions incmin

+X

+YTM14025_040




180°

270°

Clamp Band

Payload Push-OffSprings (12 Places)

0°

AdapterCone


90°


Bolt Circle Consists of0.25 in (24 0.64 cm)Holes Equally Spaced,Starting at 0º

Dimensions in cmin

Payload Attachment Plane

Side View from 0˚

Separation

Plane

Bolt Circle

3.751.48

Adapter Cone

Payload SeparationClamp Band

4.7 lbm (2.1 Kg)Remains with Payload


7.492.95

43.217.0φ

MinotaurCoordinates

+X

+YTM14025_041

+Z

+Y




Plug Shell

Receptacle Shell

Socket Contacts

Pin Contacts

SP

P S

S P

Separation Plane

Mate #1 Performedat Orbital During

Separation System Assembly

Mate #2Performed

at VAB

Plug withPin Contacts

Receptacle withSocket Contacts

PayloadInterface

Plane

Spacecraft

HarnessLengthSpecified byPayload

Can Be Suppliedto Payload

RecommendHard Mount

Launch Vehicle

Note: Sep System and Pigtails Delivered to VAB as a Unit

TM14025_043

TM14025_042Payload Weight

.5

.75

1.00

1.25

1.5

1.75

Sep

arat

ion

Vel

ocity

(m

/sec

)

Sep

arat

ion

Vel

ocity

(ft/

sec)

600

2.00

3.00

4.00

5.00

1,2001,0008006004002000 lbm

kg5004003002001000

38 in (97 cm) Interface



Figure 5-7. Payload Separation Velocities Using the Standard Separation System

Figure 5-8. Vehicle/Spacecraft Electrical Connectors and Associated Electrical Harnesses



5.3.4. Pyrotechnic Initiation SignalsOrbital provides the capability to directly

initiate 16 separate pyrotechnic conductorsthrough two dedicated Ordnance Driver Mod-ules (ODM). The ODM provides a 10 A, 100ms, current limited pulse into a 1 ±0.1 W initia-tion device.

5.3.5. Payload TelemetryStandard Minotaur service provides a

number of dedicated payload discrete (bi-level)and analog telemetry monitors through dedi-cated channels in the vehicle encoder. For dis-crete monitors, the payload customer must pro-vide the 5 Vdc source and the return path. Thecurrent at the payload interface must be lessthan 10 mA. Separation breakwire monitors canbe specified if required. The number of analogchannels available for payload telemetry moni-toring is dependent on the frequency of the data.Payload telemetry requirements and signalcharacteristics will be specified in the PayloadICD and should not change once the final te-lemetry format is released at approximately L-6 months.

5.3.6. Non Standard Electrical InterfacesNon-standard services such as serial

command and telemetry interfaces can be ne-gotiated between OSP and the payload contrac-tor on a mission-by-mission basis.

5.3.7. Electrical Launch Support EquipmentOrbital will provide space for a rack of

customer supplied EGSE in the LCR, or either ofthe on-pad equipment vaults. The equipmentwill interface with the launch vehicle/space-craft through either the dedicated payload um-bilical interface or directly through the payloadaccess door. The payload customer is respon-sible for providing cabling from the EGSE loca-tion to the launch vehicle/spacecraft.

5.4. Payload Design ConstraintsThe following sections provide design

constraints to ensure payload compatibility withthe Minotaur system.

5.4.1. Payload Center of Mass ConstraintsAlong the Y and Z axes, the payload c.g.

must be within 1.5 in (3.8 cm) of the vehicle

centerline for the standard configuration (withinthe accuracy listed in Figure 5-9). Payloadswhose c.g. extend beyond the 1.5 in. lateraloffset limit will require Orbital to verify the spe-cific offsets that can be accommodated.

Figure 5-9. Payload Mass PropertiesMeasurement Tolerance

5.4.2. Final Mass Properties AccuracyThe final mass properties statement must

specify payload weight to an accuracy of atleast 1 lbm (0.5 kg), the center of gravity to anaccuracy of at least 0.25 in (6.4 mm) in eachaxis, and the products of inertia to an accuracyof at least 0.5 slug-ft2 (0.7 kg-m2). In addition, ifthe payload uses liquid propellant, the slosh fre-quency must be provided to an accuracy of 0.2Hz, along with a summary of the method usedto determine slosh frequency.

5.4.3. Pre-Launch Electrical ConstraintsPrior to launch, all payload electrical in-

terface circuits are constrained to ensure there isno current flow greater than 10 mA across thepayload electrical interface plane. The primarysupport structure of the spacecraft shall be elec-trically conductive to establish a single point elec-trical ground.

5.4.4. Payload EMI/EMC ConstraintsThe Minotaur avionics share the payload

area inside the fairing such that radiated emis-sions compatibility is paramount. OSP placesno firm radiated emissions limits on the payloadother than the prohibition against RF transmis-sions within the payload fairing. Prior to launch,Orbital requires review of the payload radiatedemission levels (MIL-STD-461, RE02) to verifyoverall launch vehicle EMI safety margin (emis-sion) in accordance with MIL-E-6051. Payload

Measurement

Mass

Principal Moments of Inertia

Cross Products of Inertia

Center of Gravity X, Y and Z Axes

Accuracy

±1 lbm (±0.5 kg)

±5%

±0.5 sl - ft2(±0.7 kg - m2)

±6.4 mm (±0.25 in)

TM14025_045



RF transmissions are not permitted after fairingmate and prior to an ICD specified time afterseparation of the payload. An EMI/EMC analy-sis may be required to ensure RF compatibility.

Payload RF transmission frequencies mustbe coordinated with Orbital and range officialsto ensure non-interference with Minotaur andrange transmissions. Additionally, the customermust schedule all RF tests at the integration sitewith Orbital in order to obtain proper range clear-ances and protection.

5.4.5. Payload Dynamic FrequenciesTo avoid dynamic coupling of the payload

modes with the natural frequency of the vehicle,the spacecraft should be designed with a struc-tural stiffness to ensure that the lateral fundamen-tal frequency of the spacecraft, fixed at the space-craft interface is typically greater than 12 Hz.However, this value is effected significantly byother factors such as incorporation of a space-craft isolation system and/or separation system.Therefore, the final determination of compatibil-ity must be made on a mission-specific basis.

5.4.6. Payload Propellent SloshA slosh model should be provided to Or-

bital in either the pendulum or spring-mass for-mat. Data on first sloshing mode are requiredand data on higher order modes are desirable.The slosh model should be provided with thepayload finite element model submittals.

5.4.7. Payload-Supplied Separation SystemsIf the payload employs a non-Orbital sepa-

ration system, then the shock delivered to theStage 4 vehicle interface must not exceed the limitlevel characterized in Section 4.3 (Figure 4-6).As stated in that section, shock above this levelcould require a requalification of components oran acceptance of risk by the payload customer.

5.4.8. System Safety ConstraintsOSP considers the safety of personnel

and equipment to be of paramount importance.EWR 127-1 outlines the safety design criteria

for Minotaur payloads. These are compliancedocuments and must be strictly followed. It isthe responsibility of the customer to ensure thatthe payload meets all OSP, Orbital, and rangeimposed safety standards.

Customers designing payloads that employhazardous subsystems are advised to contact OSPearly in the design process to verify compliancewith system safety standards.


Minotaur Payload User's Guide Section 6.0 - Mission Integration

6. MISSION INTEGRATION

6.1. Mission Management ApproachThe Minotaur program is managed

through the US Air Force, Space and Missile Sys-tems Center, Rocket Systems Launch Program(SMC Det 12/RP). SMC Det 12/RP serves as theprimary point of contact for the payload custom-ers for the Minotaur launch service. A typicalintegrated OSP organizational structure is shownin Figure 6-1. Open communication betweenSMC Det 12/RP, Orbital, and the customer, em-phasizing timely transfer of data and prudentdecision-making, ensures efficient launch ve-hicle/payload integration operations.

6.1.1. SMC Det 12/RP Mission ReponsibilitiesThe program office for all OSP missions is

the SMC Det 12/RP. They are the primary Pointof Contact (POC) for all contractual and techni-cal coordination. SMC Det 12/RP contracts withOrbital to provide the Launch Vehicle and sepa-rately with commercial Spaceports and/or Gov-ernment Launch Ranges for launch site facilitiesand services. Once a mission is identified, SMCDet 12/RP will assign a Mission Manager to co-

ordinate all mission planning and contractingactivities. SMC Det 12/RP is supported by TRWand other associate contractors for technical andlogistical support, particularly utilizing their ex-tensive expertise and background knowledge ofthe Minuteman booster and subsystems.

6.1.2. Orbital Mission ResponsibilitiesAs the launch vehicle provider, Orbital’s

responsibilities fall into four primary areas:a. Launch Vehicle Program Managementb. Mission Managementc. Engineeringd. Launch Site Operations

Orbital assigns a Mission Manager to man-age the launch vehicle technical and program-matic interfaces for a particular mission. The Or-bital Mission Manager is the single POC for allaspects of a specific mission. This person hasoverall program authority and responsibility toensure that payload requirements are met andthat the appropriate launch vehicle services areprovided. The Orbital Mission Manager willjointly chair the Mission Integration WorkingGroups (MIWGs) with the SMC Det 12/RP Mis-

Figure 6-1. OSP Management Structure

TM14025_073

SMC Det 12/RPRocket Systems LaunchProgram (RSLP) Office

Orbital Sciences CorpOSP Prime Contractor

• Vehicle Development• Mission and Payload

Integration• Launch Operations

SMC Det 9RSLP VAFB Support

• Range Interface• Facility Control• Launch Support

TRWRSLP SETA Support

• Technical Support• Independent Analysis• MM Expertise

Ogden ALCMM Motor Processing

• MM Motor Processing• Logistics• X-Rays

Spaceports andGovernment Ranges

• Launch Facilities• Range Instrumentation• Flight Safety



sion Manager. The Mission Managers respon-sibilities include detailed mission planning, pay-load integration services, systems engineering,mission-peculiar design and analyses coordina-tion, payload interface definition, launch rangecoordination, integrated scheduling, launch siteprocessing, and flight operations.

6.2. Mission Planning and DevelopmentOSP will assist the customer with mission

planning and development associated withMinotaur launch vehicle systems. These servicesinclude interface design and configuration con-trol, development of integration processes,launch vehicle analyses, facilities planning,launch campaign planning to include range ser-vices and special operations, and integratedschedules.

The procurement, analysis, integration andtest activities required to place a customer’s pay-load into orbit are typically conducted over a 20month long standard sequence of events calledthe Mission Cycle. This cycle normally begins18 months before launch, and extends to eightweeks after launch.

Once contract authority to proceed is re-ceived, the Mission Cycle is initiated. The con-tract option designates the payload, launch date,and basic mission parameters. In response, theMinotaur Program Manager designates an Or-bital Mission Manager who ensures that thelaunch service is supplied efficiently, reliably,and on-schedule.

The typical Mission Cycle interweaves thefollowing activities:

a. Mission management, document ex-changes, meetings, and formal re-views required to coordinate andmanage the launch service.

b. Mission analyses and payload inte-gration, document exchanges, andmeetings.

c. Design, review, procurement, testingand integration of all mission-pecu-liar hardware and software.

d. Range interface, safety, and flightoperations activities, document ex-changes, meetings and reviews.

Figure 6-2 details the typical Mission Cyclefor a specific launch and how this cycle foldsinto the Orbital vehicle production schedule withtypical payload activities and milestones. A typi-cal Mission Cycle is based on a 18 month inter-val between mission authorization and launch.This interval reflects the OSP contractual sched-ule and has been shown to be an efficient sched-ule based on Orbital’s Taurus and Pegasus pro-gram experience. However, OSP is flexible tonegotiate either accelerated cycles, which takeadvantage of the Minotaur/Pegasus multi-cus-tomer production sets, or extended cycles re-quired by unusual payload requirements, suchas extensive analysis or complex payload-launchvehicle integrated designs or tests or funding limi-tations.

6.3. Mission Integration Process

6.3.1 Integration MeetingsThe core of the mission integration pro-

cess consists of a series of Mission Integrationand Range 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-Minotaur ICD inaddition to all mission-unique analyses, hard-ware, software, and integrated procedures. TheRWG is responsible for the areas of launch siteoperations; range interfaces; safety review andapproval; and flight design, trajectory, and guid-ance. Documentation produced by the RWGincludes all required range and safety submit-tals.

Working Group membership consists ofthe Mission Manager and representatives fromMinotaur engineering and operations organiza-tions, as well as their counterparts from the cus-tomer organization. While the number of meet-

Release 1.0

March 2002

6-3

Minotaur Payload U

ser's Guide

Section 6.0 - Mission Integration

Task NameMISSION INTEGRATION

ATP

MISSION KICK-OFF

LAUNCH SITE SELECTION

MISSION DESIGN REVIEWS

ILC

PAYLOAD MILESTONES AND DOCUMENTS

MILESTONES (TYPICAL)

P/L DRAWINGS

P/L FINITE ELEMENT MODEL

P/L MASS PROPERTIES

INTEGRATION DOCUMENTATION

P/L ICD VERIFICATION

MISSION DESIGN AND ENGINEERING

PAYLOAD ICD

COUPLED LOADS ANALYSIS

MASS PROPS REPORT

TARGETING RPT

QUICK LOOK POST-FLIGHT REPORT

FINAL POST FLIGHT REPORT

SAFETY PROCESS

MSPSP (Payload Annex)

P/L HAZARDOUS PROCEDURES

RANGE SAFETY TRAJECTORY

RANGE DOCUMENTATION (UDS)

PI/SOC (AS REQ'D)

PRD/PSP(AS REQ'D)

PL PRD Annex

OR/OD

MISSION CONSTRAINTS/COUNTDOWN

LAUNCH VEHICLE MANUFACTURE

PROCURE & FAB COMPONENTS

INTEGRATION & TEST

FIELD OPERATIONS

PAYLOAD INTEGRATION

LAUNCH OPERATIONS

Mission Integration Working Groups/ Operations Working GroupsATP

MDR-1 MDR-2 MRR LRR

ILC

PDR CDR P/L TO VAB

PRELIM FINAL

PRELIM FINAL

PRELIM UPDATE FINAL (MEASURED)

TEST PLAN INTEGRATION PROCEDURES

PRELIM FINAL

PRELIM FINAL

PRELIM REV FINAL PRE-LAUNCH

PRELIM FINAL

QUICK-LOOK

FINAL

PRELIM FINAL

PRELIM FINAL

PRELIM FINAL

PI SOC

PRD PSP

PL PRD Annex

OR OD

PRELIM FINAL

M1 M2 M3 M4 M5 M6 M7 M8 M9 M10 M11 M12 M13 M14 M15 M16 M17 M18 M19 M20 M21

TM14025_074

Figure 6-2. Typical Minotaur Mission Integration Schedule



ings, both formal and informal, required to de-velop and implement the mission integration pro-cess will vary with the complexity of the space-craft, quarterly meetings are typical.

6.3.2. Mission Design Reviews (MDR)Two mission-specific design reviews will

be held to determine the status and adequacy ofthe launch vehicle mission preparations. Theyare designated MDR-1 and MDR-2 and are typi-cally held 6 months and 13 months, respectively,after authority to proceed. They are each analo-gous to Preliminary Design Reviews (PDRs) andCritical Design Reviews (CDRs), but focus pri-marily on mission-specific elements of the launchvehicle effort.

6.3.3. Readiness ReviewsDuring the integration process, reviews are

held to provide the coordination of mission par-ticipants and management outside of the regularcontact of the Working Groups. Due to the vari-ability in complexity of different payloads andmissions, the content and number of these re-views can be tailored to customer requirements.As a baseline, Orbital will conduct two readi-ness reviews as described below.

Mission Readiness Review — Conductedwithin one month of launch, the Mission Readi-ness Review (MRR) provides a pre-launch assess-ment of integrated launch vehicle/payload/facil-ity readiness prior to committing significant re-sources to 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.

6.4. DocumentationIntegration of the payload requires de-

tailed, complete, and timely preparation and sub-mittal of interface documentation. As the launchservice provider, SMC Det 12/RP is the primarycommunication path with support agencies,

which include—but are not limited to—the vari-ous Range support agencies and U.S. Govern-ment agencies such as the U.S. Department ofTransportation and U.S. State Department. Cus-tomer-provided documents represent the formalcommunication of requirements, safety data, sys-tem descriptions, and mission operations plan-ning. The major products and submittal timesassociated with these organizations are dividedinto two areas—those products that are providedby the customer, and those produced by Or-bital.

6.4.1. Customer-Provided DocumentationDocumentation produced by the cus-

tomer is detailed in the following paragraphs.

6.4.1.1. Payload QuestionnaireThe Payload Questionnaire is designed

to provide the initial definition of payload re-quirements, interface details, launch site facili-ties, and preliminary safety data to OSP. Thecustomer shall provide a response to the Pay-load Questionnaire form (Appendix A), or pro-vide the same information in a different format,in time to support the Mission Kickoff Meeting.The customer’s responses to the payload ques-tionnaire define the most current payload re-quirements and interfaces and are instrumentalin Orbital’s preparation of numerous documentsincluding the ICD, Preliminary Mission Analy-sis, and launch range documentation. Addi-tional pertinent information, as well as prelimi-nary payload drawings, should also be includedwith the response. Orbital understands that adefinitive response to some questions may notbe feasible. These items are defined during thenormal mission integration process.

6.4.1.2. Payload Mass PropertiesPayload mass properties must be pro-

vided in a timely manner in order to supportefficient launch vehicle trajectory developmentand dynamic analyses. Preliminary mass prop-erties should be submitted as part of the MRDat launch vehicle authority to proceed. Up-dated mass properties shall be provided at pre-



defined intervals identified during the initialmission integration process. Typical timing ofthese deliveries is included in Figure 6-2.

6.4.1.3. Payload Finite Element ModelA payload mathematical model is required

for use in Orbital’s preliminary coupled loadsanalyses. Acceptable forms include either aCraig-Bampton model valid to 120 Hz or aNASTRAN finite element model. For the finalcoupled loads analysis, a test verified mathemati-cal model is desired.

6.4.1.4. Payload Thermal Model for Inte-grated Thermal Analysis

An integrated thermal analysis can be per-formed for any payload as a non-standard ser-vice. A payload thermal model will be requiredfrom the payload organization for use in Orbital’sintegrated thermal analysis if it is required. Theanalysis is conducted for three mission phases:

a. Prelaunch ground operations;b. Ascent from lift-off until fairing jetti-

son; andc. Fairing jettison through payload de-

ployment.

Models must be provided in SINDA for-mat. There is no limit on model size althoughturn-around time may be increased for large mod-els.

6.4.1.5. Payload DrawingsOrbital prefers electronic versions of pay-

load configuration drawings to be used in themission specific interface control drawing, ifpossible. Orbital will work with the customer todefine the content and desired format for thedrawings.

6.4.1.6. Program Requirements Document(PRD) Mission Specific Annex Inputs

To obtain range support, a PRD must beprepared. A Minotaur PRD has been submittedand approved by the Western Range. For

launches from other Ranges, a Range-specificPRD will be created. This document describesrequirements needed to generally support theMinotaur launch vehicle. For each launch, anannex is submitted to specify the range supportneeded to meet the mission’s requirements. Thisannex includes all payload requirements as wellas any additional Minotaur requirements that mayarise to support a particular mission. The cus-tomer completes all appropriate PRD forms forsubmittal to Orbital.

6.4.1.6.1. Launch Operations Requirements(OR) Inputs

To obtain range support for the launch op-eration and associated rehearsals, an OR mustbe prepared. The customer must provide all pay-load pre-launch and launch day requirements forincorporation into the mission OR.

6.5. Safety

6.5.1. System Safety RequirementsIn the initial phases of the mission inte-

gration effort, regulations and instructions thatapply to spacecraft design and processing arereviewed. Not all safety regulations will apply toa particular mission integration activity. Tailor-ing the range requirements to the mission uniqueactivities will be the first step in establishing thesafety plan. OSP has three distinctly differentmission approaches affecting the establishmentof the safety requirements:

a. Baseline mission: Payload integrationand launch operations are conductedat VAFB, CA

b. Campaign/VAFB Payload Integrationmission: Payload integration is con-ducted at VAFB and launch opera-tions are conducted from a non-VAFBlaunch location.

c. Campaign/Non-VAFB Payload Inte-gration mission: Payload integrationand launch operations are conductedat a site other than VAFB.



For the baseline and VAFB Payload Inte-gration missions, spacecraft prelaunch operationsare conducted at Orbital’s VAB, Building 1555,VAFB. For campaign style missions, the space-craft prelaunch operations are performed at thedesired launch site.

Before a spacecraft arrives at the process-ing site, the payload organization must providethe cognizant range safety office with certifica-tion that the system has been designed and testedin accordance with applicable safety require-ments (e.g. EWR 127-1 Range Safety Require-ments for baseline and VAFB Payload Integra-tion missions). Spacecraft that integrate and/orlaunch at a site different than the processing sitemust also comply with the specific launch site’ssafety requirements. Orbital will provide the cus-tomer coordination and guidance regarding ap-plicable safety requirements.

It cannot be overstressed that the appli-cable safety requirements should be consideredin the earliest stages of spacecraft design. Pro-cessing and launch site ranges discourage the useof waivers and variances. Furthermore, approvalof such waivers cannot be guaranteed.

6.5.2. System Safety DocumentationFor each Minotaur mission, OSP acts as

the interface between the mission and RangeSafety. In order to fulfill this role, OSP requiressafety information from the payloader. Forlaunches from either the Eastern or WesternRanges, EWR 127-1 provides detailed rangesafety regulations. To obtain approval to use thelaunch site facilities, specified data must be pre-pared and submitted to the OSP Program Office.This information includes a description of eachpayload hazardous system and evidence of com-pliance with safety requirements for each sys-tem. Drawings, schematics, and assembly andhandling procedures, including proof test datafor all lifting equipment, as well as any other in-formation that will aid in assessing the respec-tive systems should be included. Major catego-ries of hazardous systems are ordnance devices,

radioactive materials, propellants, pressurizedsystems, toxic materials, cryogenics, and RF ra-diation. Procedures relating to these systems aswell as any procedures relating to lifting opera-tions or battery operations should be preparedfor safety review submittal. OSP will provide thisinformation to the appropriate safety offices forapproval.


Minotaur Payload User's Guide Section 7.0 - Ground and Launch Ops

7. GROUND AND LAUNCH OPERATIONS

7.1. Minotaur/Payload Integration OverviewThe processing of the Minotaur upper

stack utilizes many of the same proven tech-niques developed for the Pegasus and Tauruslaunch vehicles. This minimizes the handlingcomplexity for both vehicle and payload. Hori-zontal integration of the Minotaur vehicle upperstages simplifies integration procedures, increasessafety and provides excellent access for the inte-gration team. In addition, simple mechanical andelectrical interfaces reduce vehicle/payload inte-gration times, increase system reliability and mini-mize vehicle demands on payload availability.

7.2. Ground And Launch OperationsGround and launch operations are con-

ducted in three major phases:a. Launch Vehicle Integration — As-

sembly and test of the Minotaur ve-hicle

b. Payload Processing/Integration —Receipt and checkout of the satellitepayload, followed by integration withMinotaur and verification of inter-faces

c. Launch Operations — Includes trans-port of the upper stack to the launchpad, final integration, checkout, arm-ing and launch.

7.2.1. Launch Vehicle Integration

7.2.1.1. Planning and DocumentationMinotaur integration and test activities are

controlled by a comprehensive set of Work Pack-ages (WPs), that describe and document everyaspect of integrating and testing Minotaur andits payload. Mission-specific work packages arecreated for mission-unique or payload-specificprocedures. All procedures are issued byOrbital’s Vandenberg Planning Department be-fore work is started. Any discrepancies encoun-tered are recorded on a Discrepancy Report anddispositioned as required. All activities are inaccordance with Orbital’s ISO 9001 certification.

7.2.1.2. Vehicle Integration and Test Activi-ties

The major vehicle components and sub-assemblies that comprise the Minotaur UpperStack Assembly, including the Stage 3 and Stage4 Orion motors, are delivered to Orbital’s VABlocated at VAFB, CA. There, the vehicle is hori-zontally integrated prior to the arrival of the pay-load. Integration is performed at a convenientworking height, which allows relatively easy ac-cess for component installation, inspection andtest.

The integration and test process ensuresthat all vehicle components and subsystems arethoroughly tested. Since the Minuteman motorsare not available at the VAB, a high fidelity simu-lator consisting of actual Minuteman componentsis used.

7.2.1.2.1. Flight Simulation TestsFlight Simulation Tests use the actual flight

software and simulate a “fly to orbit” scenariousing simulated Inertial Navigation System (INS)data. The Flight Simulation is repeated after eachmajor change in vehicle configuration (i.e., FlightSimulation #2 after stage mate, Flight Simulation#3 with payload electrically connected (if re-quired) and Flight Simulation #4 after the pay-load is mechanically integrated). After each test,a complete review of the data is undertaken priorto proceeding. The payload nominally partici-pates in Flight Simulation #3 and #4.

7.2.2. Payload Processing/IntegrationPayloads normally undergo initial check-

out and processing at Air Force or commercialfacilities at VAFB. The payload is then sent tothe VAB for integration with the Minotaur upperstack. After arrival at the VAB, the payload com-pletes its own independent verification andcheckout prior to beginning integrated process-ing with Minotaur. Following completion ofMinotaur and payload testing the payload willbe enclosed inside the fairing. The required pay-load environments will be maintained inside the



fairing until launch. Any payload specific haz-ardous procedures should be coordinatedthrough Orbital to the launch range no later than120 days prior to first use (draft) and 30 days priorto first use (final).

7.2.2.1. Payload to Minotaur IntegrationThe integrated launch processing activi-

ties are designed to simplify final launch pro-cessing while providing a comprehensive veri-fication of the payload interface. The systemsintegration and test sequence is engineered toensure all interfaces are verified.

7.2.2.2. Pre-Mate Interface TestingIf required, the electrical interface be-

tween Minotaur and the payload is verifiedusing a mission unique Interface VerificationTest (IVT) to jointly verify that the proper func-tion of the electrical connections and com-mands. These tests, customized for each mis-sion, typically check bonding, electrical com-patibility, communications, discrete com-mands and any off nominal modes of the pay-load. After completing the IVT, a Flight Simu-lation (Flight Sim #3) is performed with thepayload electrically - but not mechanically -connected to Minotaur to demonstrate the fullsequence of events in a simulated flight sce-nario. Once Flight Sim #3 is successfully com-pleted, the payload is mechanically mated tothe launch vehicle. For payloads with simpli-fied or no electrical interfaces to Minotaur, itmay be acceptable to proceed to payload mateimmediately after the IVT. For pre-mate veri-fication of the mechanical interface, the sepa-ration system can also be made available be-fore final payload preparations.

7.2.2.3. Payload Mating and VerificationFollowing the completion of Flight Sim #3,

the jumpers between the payload and Minotaurare removed. Once the payload aft end close-outs are completed, the payload will be both me-chanically and electrically mated to the Minotaur.Following mate, the flight vehicle is ready for thefinal integrated systems test, Flight Simulation #4.

7.2.2.4. Final Processing and Fairing CloseoutAfter successful completion of Flight Simu-

lation #4, all consumables are topped off andordnance is connected. Similar payload opera-tions may occur at this time. Once consumablesare topped off, final vehicle / payload closeout isperformed and the fairing is installed. The pay-load will coordinate with OSP access to the pay-load from payload mate until final closeout be-fore launch.

7.2.2.5. Payload Propellant LoadingPayloads utilizing integral propulsion sys-

tems with propellants such as hydrazine can beloaded and secured through coordinated Orbitaland contractor arrangements for use of the pro-pellant loading facilities in the VAB. This is anon-standard service.

7.2.2.6. Final Vehicle Integration and TestDue to operational constraints, the Lower

Stack Assembly, consisting of the Minuteman mo-tors, is processed by the Air Force at a separatefacility. After testing by Orbital, it is delivereddirectly to the launch pad to await the arrival ofthe upper stack. After the vehicle is fully stackedat the pad, final tests are completed to verify ve-hicle integrity and all interfaces to the range areexercised.

7.3 Launch Operations

7.3.1. Launch Control OrganizationThe Launch Control Organization is split

into two groups: the Management group and theTechnical group. The Management group con-sists of senior range personnel and Mission Di-rectors/Managers for the launch vehicle and pay-load. The Technical Group consists of the per-sonnel responsible for the execution of the launchoperation and data review/assessment for thePayload, the Launch Vehicle and the Range. ThePayload’s members of the technical group areengineers who provide technical representationin the control center. The Launch Vehicle’smembers of the technical group are engineerswho prepare the Minotaur for flight, review and



assess data that is displayed in the Launch Con-trol Room (LCR) and provide technical represen-tation in the LCR and in the Launch OperationsControl Center (LOCC). The Range’s membersof the technical group are personnel that main-tain and monitor the voice and data equipment ,tracking facilities and all assets involved with RFcommunications with the launch vehicle. Inaddition, the Range provides personnel respon-sible for the Flight Termination System monitor-ing and commanding.


Section 8.0 - Optional EnhancedMinotaur Payload User's Guide Capabilities

8. OPTIONAL ENHANCED CAPABILITIESThe OSP launch service is structured to

provide a baseline vehicle configuration whichis then augmented with optional enhancementsto meet the unique needs of individual payloads.The baseline vehicle capabilities are defined inthe previous sections and the optional enhancedcapabilities are defined below.

8.1. Mechanical Interface and Separation Sys-tem Enhancements

8.1.1. Separation SystemsOrbital offers several alternate separating

interfaces as optional services. The baseline op-tional separation system is the Orbital 38 in. de-sign, but several flight-proven, Orbital-developedPayload Attach Fittings (PAFs) can be utilized.These systems are discussed in Section 5.2.1 and5.2.2. In addition, other separation systems rang-ing from an Orbital-developed 10 in. diametersystem to the SAAB 38.81 in. system can be pro-vided as additional negotiated options.

8.1.2. Additional Fairing Access DoorsOSP provides one 8.5 in x 13.0 in (21.6

cm x 33.0 cm), graphite, RF-opaque payload fair-ing access door (per Section 5.1.2). Additionaldoors can be provided within the range speci-fied in Section 5.1.2. Other fairing access con-figurations, such as small circular access panels,can also be provided as negotiated mission-spe-cific enhancements.

8.1.3. Payload Isolation SystemOSP offers a flight-proven payload isola-

tion system as a non-standard service. ThisSoftride for Small Satellites (SRSS) was developedby Air Force Research Laboratory (AFRL) and CSAEngineering. It was successfully demonstratedon the two initial Minotaur missions in the typi-cal configuration shown in Figure 8-1. This me-chanical isolation system has demonstrated thecapability to significantly alleviate the transientdynamic loads that occur during flight. The iso-lation system can provide relief to both the over-all payload center of gravity loads and compo-

nent or subsystem responses. Typically the sys-tem will reduce transient loads to approximately50% of the level they would be without the sys-tem. The exact results can be expected to varyfor each particular spacecraft and with locationon the spacecraft. Generally, a beneficial reduc-tion in shock and vibration will also be provided.The isolation system does impact overall vehicleperformance (by approximately 20-40 lb [9-18kg]) and the available payload dynamic enve-lope by up to 4.0 in (10.16 cm) axially and up to1.0 in (2.54 cm) laterally.

8.1.4. Increased Payload VolumeTo accommodate payloads larger than

those that can be accommodated by the stan-

Figure 8-1. Softride for Small Satellites (SRSS)Payload Isolation System



dard Pegasus-based fairing, a larger fairing de-sign is being developed in conjunction withAFRL. A preliminary dynamic envelope isshown in Figure 8-2. A slight performance de-crease will be incurred.

8.2. Performance Enhancements

8.2.1. Insertion AccuracyEnhanced insertion accuracy or support

for multiple payload insertion can be providedas an enhanced option utilizing the HydrazineAuxiliary Propulsion System (HAPS). Orbital in-sertion accuracy can typically be improved to±10 nm (±18.5 km) or better in each apse and±0.1 deg in inclination. For orbits above 600km, the HAPS can also increase payload mass

175.0

162.3

133.5

82.2

0.0

54.7

35.6

INTERFACE PLANE

22.7

AVAILABLEENVELOPE(DYNAMIC)

TM14025_075

Figure 8-2. Optional 61 in. Diameter Fairing

by approximately 50 to 250 lbm, depending onthe orbit. Specific performance capability as-sociated with the HAPS can be provided by con-tacting the OSP program office. HAPS is moreeffective at higher altitude and also permits in-jection of shared payloads into different orbits.HAPS, which is mounted inside the AvionicsStructure, consists of a hydrazine propulsionsubsystem and a Stage 4 separation subsystem.After burnout and separation from the Stage 4motor, the HAPS hydrazine thrusters provideadditional velocity and both improved perfor-mance and precise orbit injection. The HAPSpropulsion subsystem consists of a centrallymounted tank containing approximately 130lbm (59 kg) of hydrazine, helium pressurizationgas, and three fixed, axially pointed thrusters.The hydrazine tank contains an integral blad-der which will support multiple restarts.

8.3. Environmental Control Options

8.3.1. Conditioned Air

Conditioned air can be provided withinthe fairing volume using an Environmental Con-trol System (ECS) via a duct that is retracted inthe last minutes prior to launch. Temperatureand humidity can be regulated within the limitsindicated in Figure 4-10. A filter is installed toprovide a Class 100,000 environment, typically.Upon exercise of the Enhanced Cleanliness op-tion, a certified HEPA filter is used in the inputduct to assure the necessary low particulate en-vironment (Class 100,000 or Class 10,000).

8.3.2. Nitrogen PurgeOSP can provide gaseous nitrogen purges

to the payload after fairing encapsulation untillift-off. The system distribution lines are routedalong the inner surface of the fairing. If requiredfor spot cooling of a payload component, Or-bital will provide GN2 flow to localized regionsin the fairing. The GN2 will meet Grade B speci-fications, as defined in MIL-P-27401C and canbe regulated between 2.4-11.8 l/sec (5-25 scfm).The GN2 is on/off controllable in the launchequipment vault and in the launch control room.



The system’s regulators are set to a de-sired flow rate during prelaunch processing.The system cannot be adjusted after the launchpad has been cleared of personnel.

Payload purge requirements must be co-ordinated with Orbital via the ICD to ensure thatthe requirements can be achieved.

8.3.3. Enhanced Contamination ControlUnderstanding that some payloads have

requirements for enhanced cleanliness, OSP of-fers a contamination control option, which iscomposed of the elements in the following sec-tions (which is also discussed in Section 4.7).Minotaur customers can also coordinate combi-nations of the elements listed below to meet theunique needs of their payloads.

8.3.3.1 High Cleanliness Integration Environ-ment (Class 10K or 100K)

With enhanced contamination control, asoft walled clean room can be provided to en-sure a FED-STD-209 Class M6.5 (100,000) orClass M5.5 (10,000) environment during all pay-load processing activities up to fairing encapsu-lation. The soft walled clean room andanteroom(s) utilize HEPA filter units to filter theair and hydrocarbon content is maintained at 15ppm or less. The payload organization is respon-sible for providing the necessary clean room gar-ments for payload staff as well as vehicle staffthat need to work inside the clean room.

8.3.3.2 Fairing Surface Cleanliness OptionsThe inner surface of the fairing and pay-

load cone assemblies can be cleaned to cleanli-ness criteria which ensures no particulate mattervisible with normal vision when inspected from6 to 18 inches under 100 ft-candle incident light.The same will be true when the surface is illumi-nated using black light, 3200 to 3800 Angstroms(Visibly Clean Plus Ultraviolet). In addition, Or-bital can ensure that all materials used withinthe encapsulated volume have outgassing

characteristics of less than 1.0% TML and lessthan 0.1% CVCM. Items that don’t meet theselevels can be masked to ensure they are encap-sulated and will have no significant effect on thepayload.

8.3.3.3 High Cleanliness Fairing EnvironmentWith the enhanced contamination control

option, Orbital provides an ECS from payloadencapsulation until just prior to vehicle lift-off.The ECS continuously purges the fairing volumewith clean filtered air. Orbital’s ECS incorpo-rates a HEPA filter unit to provide FED-STD-209Class M5.5 (10,000) air. Orbital monitors thesupply air for particulate matter via a probe in-stalled upstream of the fairing inlet duct prior toconnecting the air source to the payload fairing.

8.4. Enhanced Telemetry OptionsOSP can provide mission specific instru-

mentation and telemetry components to supportadditional payload or experiment data acquisi-tion requirements. Telemetry options include ad-ditional payload-dedicated bandwidth and GPS-based precision navigation data.

8.4.1. Enhanced Telemetry BandwidthA second telemetry data stream capable

of up to 2 Mbps data rate can be provided. Maxi-mum data rates depend on the mission cover-age required and the launch range receivercharacteristics and configuration. This capa-bility was successfully demonstrated on the in-augural Minotaur mission.

8.4.2. Enhanced Telemetry InstrumentationTo support the higher data rate capability

in Section 8.4.1, enhanced telemetry instrumen-tation can be provided. The instrumentation caninclude strain gauges, temperature sensors, ac-celerometers, analog data, and digital data con-figured to mission-specific requirements. Thiscapability was successfully demonstrated on theinaugural Minotaur mission.



8.4.3. Navigation DataPrecision navigation data using an inde-

pendent GPS-receiver and telemetry link is avail-able as an enhanced option. This option utilizesthe Orbital-developed GPS Position Beacon(GPB) and provides a better than 100 m positionaccuracy with a nominal 1 Hz data rate. Thiscapability was successfully demonstrated on theinaugural Minotaur mission.


Section 9.0 - Shared LaunchMinotaur Payload User's Guide Accommodations

9. SHARED LAUNCH ACCOMMODATIONSMinotaur is uniquely capable of provid-

ing launches of multiple satellite payloads, le-veraging SMC Det 12/RP and Orbital’s exten-sive experience in integrating and launchingmultiple payloads. In addition to the five satel-lites successfully separated on the inauguralMinotaur JAWSAT mission, multiple spacecraftconfigurations have been flown on many ofOrbital’s Pegasus and Taurus missions to date.

Two technical approaches are availablefor accommodating multiple payloads. Thesedesign approaches are:

Load-Bearing Spacecraft — aft spacecraftdesigned to provide the structural loadpath between the forward payload and thelaunch vehicle, maximizing utilization ofavailable mass performance and payloadfairing volume

Non Load-Bearing Spacecraft — aftspacecraft whose design cannot providethe necessary structural load path for theforward payload

9.1. Load-Bearing SpacecraftProviding a load-bearing aft payload maxi-

mizes use of available volume and mass. Theavailable mass for the aft payload is determinedby the Minotaur performance capability to orbitless the forward payload and attach hardwaremass. All remaining mission performance, ex-cluding a stack margin, is available to the aftpayload. The load-bearing spacecraft interfacesdirectly to the avionics assembly interface andthe forward payload via pre-determined inter-faces. These interfaces include standard Orbitalseparation systems and pass-through electricalconnectors to service the forward payload. Fig-ure 9-1 illustrates this approach.

Two approaches may be taken for load-bearing spacecraft. The first approach is to use adesign developed by other spacecraft suppliers,which must satisfy Minotaur and forward pay-load structural design criteria. The principal re-quirements levied upon load-bearing spacecraft

are those involving mechanical and electricalcompatibility with the forward payload. Struc-tural loads from the forward payload during allflight events must be transmitted through the aftpayload to the Minotaur. Orbital will provideminimum structural interface design criteria forshear, bending moment, axial and lateral loads,and stiffness.

The second approach involves the use ofan Orbital design using the MicroStar bus, suc-cessfully developed and flown for ORBCOMMspacecraft. The MicroStar bus features a circulardesign with an innovative, low-shock separationsystem. The spacecraft bus is designed to allowstacking of co-manifested payloads in “slices”within the fairing. The bus design is compact andprovides exceptional lateral stiffness.

A variation on the load bearing spacecraftapproach was flown on the JAWSAT inauguralmission, in which the primary payload (JAWSAT)was a Multiple Payload Adapter (MPA) fromwhich four small satellites were separated (Fig-

Figure 9-1. Typical Load Bearing SpacecraftConfiguration

TypicalForward

SpacecraftVolume

TypicalAft Load Bearing

SpacecraftVolume

(2.95)

φ 46.00(Dynamic)

(3.95)

FairingDynamicEnvelope

φ 23.00 in.Separation Ring

φ 38 in. Avionics Thrust Tube(22.00 in. Long)

φ 38 in.Separation Ring

TM14025_077



ure 9-2). After separating the smaller “piggy-back” satellites, the JAWSAT MPA was alsoseparated as an autonomous satellite by utiliz-ing the Orbital 23 in. separation system andadapter cone. An updated concept to providegreater payload options and primary payloadvolume (by mounting directly to the avionicsassembly) is shown in Figure 9-3.

Integrated coupled loads analyses will beperformed with test verified math models pro-vided by the payload. These analyses are requiredto verify the fundamental frequency and deflec-tions of the stack for compliance with theMinotaur requirement of 12 Hz minimum. De-sign criteria provided by OSP will include “stack”margins to minimize interactive effects associ-ated with potential design changes of each pay-load. OSP will provide the necessary engineer-ing coordination between the spacecraft andlaunch vehicle.

Electrical pass-through harnesses will alsoneed to be provided by the aft payload alongwith provisions for connectors and interface veri-fication. The spacecraft supplier will need to pro-vide details of the appropriate analyses and teststo OSP to verify adequacy of margins and showthat there is no impact to the forward spacecraftor the launch vehicle.

Figure 9-2. JAWSAT Multiple PayloadAdapter Load Bearing Spacecraft

9.2. Non Load-Bearing SpacecraftFor aft spacecraft that are not designed for

withstanding and transmitting structural loadsfrom the forward payload, several options arepossible including AFRL and Orbital design con-cepts, as well as the flight-proven Dual PayloadAttach Fitting (DPAF).

The DPAF structure (Figure 9-4) is an allgraphite structure which provides independentload paths for each satellite. The worst-case “de-sign payload” for the DPAF is a 425 lbm (193kg) spacecraft with a 20 in. (51 cm) center ofmass offset and first lateral frequency of typically12 Hz. The DPAF is designed to accommodatethis “design payload” at both the forward and aftlocations, although the combined mass of the twopayloads cannot exceed Minotaur capabilities.The upper spacecraft loads are transmittedaround the lower spacecraft via the DPAF struc-ture, thus avoiding any structural interface be-tween the two payloads.

TM14025_078

TM14025_079

Figure 9-3. Five Bay Multiple Payload AdapterConcept



Primary Payload Volume

+Y

+X

φ 76.0 29.9

φ 66.026.0

Ogive Radius 269.2 106.0

101.640.0

102.940.5

Primary Payload Separation Plane

φ 114.345.0

DynamicEnvelope

Minotaur Avionics

Adapter Cone Separation PlaneBeginning of Ogive

Secondary Payload Separation Plane

43 Separation17 System

97 Separation 38 System

128.850.7

55.922.0

AvailableSecondary

PayloadVolume

5823

Separation System

Dimensions in cmin

TM14025_082

Figure 9-4. Dual Payload Attach Fitting Configuration

Release 1.0 March 2002 A-1

Minotaur Payload User's Guide Appendix A

Appendix APayload Questionnaire

MINOTAURPAYLOAD QUESTIONNAIRE

1

SATELLITE IDENTIFICATION

FULL NAME:

ACRONYM:

OWNER/OPERATOR:

INTEGRATOR(s):

ORBIT INSERTION REQUIREMENTS*

SPHEROID q Standard (WGS-84, Re = 6378.137 km)

q Other:

ALTITUDE Insertion Apse:

q nmi ___ ± __ q km

Opposite Apse:

q nmi ± q km

or... Semi-Major Axis:

q nmi ___ ± __ q km

Eccentricity:

≤ e ≤

INCLINATION ± deg

ORIENTATION Argument of Perigee:

± deg

Longitude of Ascending Node (LAN):

± deg

Right Ascension of Ascending Node (RAAN):

± deg ...for Launch Date: * Note: Mean orbital elements

LAUNCH WINDOW REQUIREMENTS

NOMINAL LAUNCH DATE:

OTHER CONSTRAINTS (if not already implicit from LAN or RAAN requirements, e.g., solar beta angle,eclipse time constraints, early on-orbit ops, etc):


2

EARLY ON-ORBIT OPERATIONS

Briefly describe the satellite early on-orbit operations, e.g., event triggers (separation sense, sun acquisition,etc), array deployment(s), spin ups/downs, etc:

SATELLITE SEPARATION REQUIREMENTS

ACCELERATION Longitudinal: � g’s Lateral: � g’s

VELOCITY Relative Separation Velocity Constraints:

ANGULAR RATES

(pre-separation) Longitudinal: Pitch: ± deg/sec

± deg/sec Yaw: ± deg/sec

ANGULAR RATES

(post-separation) Longitudinal: Pitch: ± deg/sec

± deg/sec Yaw: ± deg/sec

ATTITUDE

(at deployment)

Describe Pointing Requirements Including Tolerances:

SPIN UP Longitudinal Spin Rate: ± deg/sec

OTHER Describe Any Other Separation Requirements:

SPACECRAFT COORDINATE SYSTEM

Describe the Origin and Orientation of the spacecraft reference coordinate system, including its orientation withrespect to the launch vehicle (provide illustration if available):


3

SPACECRAFT PHYSICAL DIMENSIONS

STOWED

CONFIGURATION

Length/Height: Diameter:

q in q cm q in q cm

Other Pertinent Dimension(s):

Describe any appendages/antennas/etc which extend beyond the basic satellite envelope:

ON-ORBIT

CONFIGURATION

Describe size and shape:

If available, provide dimensioned drawings for both stowed and on-orbit configurations.

SPACECRAFT MASS PROPERTIES*

PRE-SEPARATION Inertia units: q lbm-in2 q kg-m2

Mass: q lbm q kg

Ixx:

Xcg: q in q cm Iyy:

Izz:

Ycg: q in q cm Ixy:

Iyz:

Zcg: q in q cm Ixz:

POST-SEPARATION

(non-separatingadapter remaining with

launch vehicle)

Inertia units: q lbm-in2 q kg-m2

Mass: q lbm q kg

Ixx:

Xcg: q in q cm Iyy:

Izz:

Ycg: q in q cm Ixy:

Iyz:

Zcg: q in q cm Ixz:

* Stowed configuration, spacecraft coordinate frame


4

ASCENT TRAJECTORY REQUIREMENTS

Free Molecular Heating at Fairing Separation:(Standard Service: � 360 Btu/ft2/hr)

q Btu/ft2/hr FMH � q W/m2

Fairing Internal Wall Temperature(Standard Service: � 200°F)

q deg F T � q deg C

Dynamic Pressure at Fairing Separation:(Standard Service: � 0.01 lbf /ft

2)q lbf /ft

2

q � q N/m2

Ambient Pressure at Fairing Separation:(Standard Service: � 0.3 psia)

q lbf /in2

P � q N/m2

Maximum Pressure Decay During Ascent:(Standard Service: � 0.6 psia)

q lbf /in2/sec

�� 3�� q N/m2/sec

Thermal Maneuvers During Coast Periods:(Standard Service: none)

SPACECRAFT ENVIRONMENTS

THERMALDISSIPATION

Spacecraft Thermal Dissipation, Pre-Launch Encapsulated: Watts

Approximate Location of Heat Source:

TEMPERATURE Temperature Limits During Max q deg F q deg C

Ground/Launch Operations: Min q deg F q deg C

(Standard Service is 55°F to 80°F)

Component(s) Driving Temperature Constraint:Approximate Location(s):

HUMIDITY Relative Humidity: or, Dew Point:

Max % Max q deg F q deg C

Min % Min q deg F q deg C

(Standard Service is 37 deg F)

NITROGENPURGE

Specify Any Nitrogen Purge Requirements, Including Component Description, Location,and Required Flow Rate:

(Nitrogen Purge is a Non-Standard Service)

CLEANLINESS Volumetric Requirements (e.g. Class 100,000): Surface Cleanliness (e.g. Visually Clean): Other:

LOAD LIMITS Ground Transportation Load Limits: Axial � g’s Lateral � g’s


5

ELECTRICAL INTERFACE

Bonding Requirements:

Are Launch Vehicle Supplied

Pyro Commands Required? Yes / No If Yes, magnitude: amps for msec

(Standard Service is 10 amps maximum for 100 msec)

Are Launch Vehicle Supplied If Yes, describe:

Discrete Commands Required? Yes / No

Is Electrical Access to the Satellite Required... After Encapsulation? Yes / No

at Launch Site Yes / No

Is Satellite Battery Charging Required... After Encapsulation? Yes / No

at Launch Site? Yes / No

Is a Telemetry Interface with the Launch Vehicle Flight Computer Required? Yes / No

If Yes, describe:

Other Electrical Requirements:

Please complete attached sheet of required pass-through signals.

RF RADIATION

Time After Separation Until RF Devices Are Activated:

(Note: Typically, no spacecraft radiation is allowed from encapsulation until 30 minutes after liftoff.)

Frequency: MHz Power: Watts

Location(s) on Satellite (spacecraft coordinate frame):

Longitudinal q in q cm Clocking (deg), Describe:



6

REQUIRED PASS-THROUGH SIGNALSItem

# Pin Signal Name From LEV To Satellite ShieldingMax

Current(amps)

Total LineResistance

(ohms)1

2

3

4

5

6

7

8

9

10

11

12

13

14

15

16

17

18

19

20

21

22

23

24

25

26

27

28


7

MECHANICAL INTERFACE

DIAMETER Describe Diameter of Interface (e.g. Bolt Circle, etc):

SEPARATION

SYSTEM

Will Launch Vehicle Supply the Separation System? Yes / No

If Yes approximate location of electrical connectors:

special thermal finishes (tape, paint, MLI) needed:

If No, provide a brief description of the proposed system:

SURFACE

FLATNESS

Flatness Requirements for Sep System or Mating Surface of Launch Vehicle:

FAIRING

ACCESS

Payload Fairing Access Doors (spacecraft coordinate frame):



Note: Standard Service is one door

DYNAMICS Spacecraft Natural Frequency:

Axial Hz Lateral Hz

Recommended: > 35 Hz > 12 Hz

OTHER Other Mechanical Interface Requirements:


8

GROUND SUPPORT EQUIPMENT

Describe any additional control facilities (other than the baseline Support Equipment Building (SEB) andLaunch Equipment Vault (LEV)) which the satellite intends to use:

SEB Describe (in the table below) Satellite EGSE to be located in the SEB.

[Note: Space limitations exist in the SEB, 350 ft umbilical cable length to spacecraft typical]

Equipment Name / Type

..................................................................

..................................................................

..................................................................

..................................................................

..................................................................

Approximate Size (LxWxH)

...............................................

...............................................

...............................................

...............................................

...............................................

Power Requirements

................................

................................

................................

................................

................................

Is UPS required for equipment in the SEB? Yes / No

Is Phone/Fax connection required in the SEB? Yes / No Circle: Phone / FAX

LEV Describe (in the table below) Satellite EGSE to be located in the LEV.

[Note: Space limitations exist in the SEB, 150 ft umbilical cable length to spacecraft typical]Equipment Name / Type

..................................................................

..................................................................

..................................................................

..................................................................

..................................................................

Approximate Size (LxWxH)

...............................................

...............................................

...............................................

...............................................

...............................................

Power Requirements

................................

................................

................................

................................

................................

Is UPS required for equipment in the LEV? Yes / No

Is Ethernet connection between SEB and LEV required? Yes / No

Release 1.0 March 2002 B-1

Minotaur Payload User's Guide Appendix B

APPENDIX BElectrical Interface Connectors

There are two electrical interface connectors on the payload separation plane. The first connector isspecifically designed for ordnance. Its part number is MS27474T14F-18SN. The second connectoris designed for power, payload battery charging, discrete commands, discrete telemetry, separationindicators, analog telemetry, and serial communication. Its part number is MS27474T16F-42SN.Typical circuits passing through the payload-to-launch vehicle electrical connections are shown inFigure B-1.

Release 1.0 March 2002 B-2

Minotaur Payload User's Guide Appendix B

Launch Vehicle

10 Amp, 100 msec

5V

Spacecraft

Note: Channels Can be Reallocated Based on Customer Requirements

Standard Quantity

5

5

3

4 Redundant Pairs

6

Separation Indicator

Pyro Initiation

Analog Telemetry20 mVpp to 10 Vpp

±50 V

Serial Communication Link

RS-422/RS-485

GSE Pass-Through Wiring 60 Conductors(Various AWG/Shielding Available)

Discrete Commands

5VDiscrete Telemetry

Monitors (5 Hz)

1

TM14025_080

Figure B-1. Typical Minotaur Payload Electrical Interface Block Diagram

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