Minotaur User's Guide · ACS Attitude Control System AFRL Air Force Research Laboratory AODS...

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

)


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)


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)


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



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

Minotaur User's Guide · ACS Attitude Control System AFRL Air Force Research Laboratory AODS...

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