Orbital Minotaur VI User Manual

June 2013 Minotaur IV • V • VI User’s Guide

Release 2.0

Approved for Public Release
Distribution Unlimited

©2013 Orbital Sciences Corporation

All Rights Reserved.

Minotaur IV • V • VI User’s Guide Revision Summary

1.0

Initial Release

General nomenclature, history, and administrative
updates (no technical updates)

1.

2.

Extensively Revised

REVISION SUMMARY

VERSION DOCUMENT DATE CHANGE PAGE

TM-17589 Jan 2005

1.1 TM-17589A Jan 2006

2.0 TM-17589B Jun 2013

All

All

Updated launch history
Corrected contact information

All

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Minotaur IV • V • VI User’s Guide Preface

PREFACE

The information provided in this user’s guide is for initial planning purposes only. Information for
development/design is d etermined through mission specific engineering analyses. The results of these
analyses are documented in a mission-specific Interface Control Document (ICD) for the payloader
organization to use in their development/design process. This document provides an overview of the
Minotaur system design and a description of the services provided to our customers.

Additional technical inf ormation and copies of this User's Guide may be requested from Orbital at:

[e-mail]
www.orbital.com/spacelaunch/Minotaur/IV

[phone]

Minotaur Program - Mission Development
Orbital Sciences Corporation
Launch Systems Group
3380 S. Price Road
Chandler, AZ 85248

Additional informat ion can be obtained from the USAF OSP Office at:

USAF SMC Space Development and Test Directorate (SMC/SD)
Launch Systems Division (SMC/SDL)
3548 Aberdeen Ave SE
Kirtland AFB, NM 87117-5778

[phone]
[phone]

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1. INTRODUCTION ................................................................................................................................. 1

1.1. Minotaur Family Performance and Capability ............................................................................... 2

2. MINOTAUR IV CONFIGURATIONS ................................................................................................... 4

2.1. Minotaur IV Launc h S ystem Overview .......................................................................................... 4

2.2. Minotaur IV Launc h Serv ic e .......................................................................................................... 5

2.3. Minotaur IV Launc h Veh icle .......................................................................................................... 5

2.3.1. Stage 1, 2 and 3 Booster Assemblies ....................................................................................... 5

2.3.2. Upper Stage Propulsion ............................................................................................................ 6

2.3.3. Guidance and Control Assembly (GCA) .................................................................................... 7

2.3.3.1. Avionics ............................................................................................................................... 8

2.3.3.2. Attitude Control Systems ..................................................................................................... 8

2.3.3.3. Telemetry Subsystem.......................................................................................................... 8

2.3.4. Payload Interface ....................................................................................................................... 9

2.3.5. Payload Fairing ........................................................................................................................ 10

2.3.6. Minotaur IV Launch Vehicle Enhanced Performance Configurations ..................................... 11

2.3.6.1. Minotaur IV+ (STAR 48 Stage 4) ...................................................................................... 11

2.3.6.2. Minotaur V (High Energy Performance) ............................................................................ 12

2.3.6.3. Minotaur VI ........................................................................................................................ 13

2.3.6.4. Minotaur VI+ (High Energ y Performance) ......................................................................... 14

2.4. Launch Support Equipment ........................................................................................................ 14

3. GENERAL PERFOR M AN C E ............................................................................................................ 17

3.1. Mission Profiles ........................................................................................................................... 17

3.2. Launch Sites ............................................................................................................................... 17

3.2.1. Western Launch Sites ............................................................................................................. 19

3.2.2. Eastern Launch Sites .............................................................................................................. 19

3.2.3. Alternate Launch Sites ............................................................................................................ 19

3.3. Performance Capability ............................................................................................................... 20

3.3.1. Minotaur IV LEO Orbits ........................................................................................................... 22

3.3.2. Minotaur IV+ LEO Orbits ......................................................................................................... 24

3.3.3. Minotaur VI LEO Orbits ........................................................................................................... 26

3.3.4. Elliptical Orbits and High Energ y Orbits .................................................................................. 28

3.4. Injection Accuracy ....................................................................................................................... 32

3.5. Payload Dep lo yment ................................................................................................................... 32

3.6. Payload Separ at ion ..................................................................................................................... 32

3.7. Collision/Contamination Avoidance Maneuver ........................................................................... 33

4. PAYLOAD ENVIRONMENT .............................................................................................................. 34

4.1. Steady State and Transient Acceleration Loads......................................................................... 34

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4.1.1. Transient Loads ....................................................................................................................... 34

4.1.2. Steady-State Acceleration ....................................................................................................... 36

4.2. Payload Vibration Environment................................................................................................... 36

4.2.1. Random Vibration .................................................................................................................... 37

4.2.2. Sine Vibration .......................................................................................................................... 37

4.3. Payload Acoustic Environment ................................................................................................... 39

4.4. Payload Shock Environment ....................................................................................................... 40

4.5. Payload Structural Integrity and Environments Verification ........................................................ 42

4.6. Thermal and Humidity Environments .......................................................................................... 42

4.6.1. Ground Operations .................................................................................................................. 42

4.6.2. Powered Flight ......................................................................................................................... 43

4.6.3. Nitrogen Purge (Non-Standard Service) ................................................................................. 44

4.7. Payload Contamination Control .................................................................................................. 44

4.8. Payload Elec tromagnetic Environment ....................................................................................... 45

5. PAYLOAD INTERFACES .................................................................................................................. 46

5.1. Payload Fairing ........................................................................................................................... 46

5.1.1. 92” Standard Minotaur Fairing ................................................................................................. 46

5.1.1.1. 92” Fairing Payload Dynamic Design Envelope ................................................................ 46

5.1.2. Optional 110” Fairing ............................................................................................................... 47

5.1.2.1. 110” Fairing Payload Dynamic Design Envelope .............................................................. 47

5.1.3. Payload Access Door .............................................................................................................. 48

5.2. Payload Mechanical Interface and Separation System .............................................................. 48

5.2.1. Minotaur Coordinate S ystem ................................................................................................... 49

5.2.2. Orbital Supplied Mechanical Interface Control Drawing .......................................................... 50

5.2.3. Standard Non-Separating Mechanical Interface ..................................................................... 50

5.2.4. Optional Mechanical Interfaces ............................................................................................... 50

5.2.4.1. Payload Cone Interfaces ................................................................................................... 52

5.2.4.2. Dual and Multi Payload Adapter Fittings ........................................................................... 52

5.2.4.2.1. Dual-Payload Adapter Fitting ...................................................................................... 52

5.2.4.2.2. Multi-Payload Adapter Fitting (MPAF) ........................................................................ 54

5.2.4.2.3. Minotaur V and VI+ Payload Adapter Fitting............................................................... 55

5.2.5. Optional Separation Systems .................................................................................................. 55

5.2.5.1. Orbital 38” Separ at ion System .......................................................................................... 57

5.2.5.2. Planetary Systems Motorized Lightband (MLB) ................................................................ 58

5.2.5.3. RUAG 937 Separation S ystems ........................................................................................ 59

5.3. Payload Electrical Interfaces....................................................................................................... 59

5.3.1. Payload Umbilical Interfaces ................................................................................................... 59

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5.3.2. Payload Interface Circuitry ...................................................................................................... 60

5.3.3. Payload Battery Charging ........................................................................................................ 60

5.3.4. Payload Command and Control .............................................................................................. 61

5.3.5. Pyrotechnic Initiation Signals ................................................................................................... 61

5.3.6. Payload Telemetry ................................................................................................................... 61

5.3.7. Payload Separation Monitor Loopbacks .................................................................................. 61

5.3.8. Telemetry Interfaces ................................................................................................................ 61

5.3.9. Non-Standard Electrical Interfaces .......................................................................................... 62

5.3.10. Electrical Launch Suppor t Equ ipment ................................................................................... 62

5.4. Payload Design Constraints........................................................................................................ 62

5.4.1. Payload Center of Mass Constraints ....................................................................................... 62

5.4.2. Final Mass Properties Accuracy .............................................................................................. 62

5.4.3. Pre-Launch Electrical Constraints ........................................................................................... 62

5.4.4. Payload EMI/EMC Constraints ................................................................................................ 62

5.4.5. Payload Dynamic Frequencies ................................................................................................ 63

5.4.6. Payload Propellant Slosh ........................................................................................................ 63

5.4.7. Payload-Supplied Separation Systems ................................................................................... 63

5.4.8. System Safety Constraints ...................................................................................................... 63

6. MISSION INTEGRATION .................................................................................................................. 64

6.1. Mission Management Approach ................................................................................................. 64

6.1.1. SD/SDL Mission Responsibilities ............................................................................................ 64

6.1.2. Orbital Mission Responsibilities ............................................................................................... 64

6.2. Mission Planni ng and D e velopment ........................................................................................... 65

6.2.1. Mission Assurance .................................................................................................................. 67

6.3. Mission Integration Process ........................................................................................................ 68

6.3.1. Integration Meetings ................................................................................................................ 68

6.3.2. Mission Design Reviews (MDR) .............................................................................................. 68

6.3.3. Readiness Reviews ................................................................................................................. 68

6.4. Documentation ............................................................................................................................ 69

6.4.1. Customer-Provided Documentation ........................................................................................ 69

6.4.1.1. Payload Questionnaire ...................................................................................................... 69

6.4.1.2. ICD Inputs ......................................................................................................................... 69

6.4.1.3. Payload Mass Properties .................................................................................................. 69

6.4.1.4. Payload Finite Element Model .......................................................................................... 69

6.4.1.5. Payload Thermal Model for Integrated Thermal Analysis ................................................. 70

6.4.1.6. Payload Drawings ............................................................................................................. 70

6.4.1.7. Program Requirements Document (PRD) Mission Specific Annex Inputs ....................... 70

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6.4.1.7.1. Launch Operations Requirements (OR) Inputs .......................................................... 70

6.4.1.8. Payload Launch Site Integration Procedures .................................................................... 70

6.4.1.9. ICD Verification Documentation ........................................................................................ 70

6.4.2. Orbital Produced Documentation, Data, and Analyses ........................................................... 70

6.4.2.1. Launch Vehicle to Payload ICD ........................................................................................ 71

6.4.2.2. ICD Verification Documentation ........................................................................................ 71

6.4.2.3. Preliminary Mission Analyses ........................................................................................... 71

6.4.2.4. Coupled Loads Analyses (CLA) ........................................................................................ 71

6.4.2.5. Integrated Launch Site Pro cedur es ................................................................................... 71

6.4.2.6. Missile System Pre-Launch Safety Package (MSPSP) Annex ......................................... 72

6.4.2.7. PRD Mission Specific Annex ............................................................................................. 72

6.4.2.8. Launch Operation Requirements (OR) ............................................................................. 72

6.4.2.9. Mission Constraints Document (MCD) .............................................................................. 72

6.4.2.10. Final Countdown Procedure ............................................................................................ 72

6.4.2.11. Post-Launch Analyses..................................................................................................... 72

6.5. Safety .......................................................................................................................................... 73

6.5.1. System Safety Requirements .................................................................................................. 73

6.5.2. System Safety Documentation ................................................................................................ 73

7. GROUND AND LAUNCH OPERATIONS ......................................................................................... 74

7.1. Launch Vehicle Integration Overview ......................................................................................... 75

7.1.1. Planning and Documentation .................................................................................................. 75

7.1.2. Guidance and Control Assembly Integration and Test Activities ............................................ 75

7.1.3. PK Motor Integration and Test Activities ................................................................................. 76

7.1.4. Mission Simulation Tests ......................................................................................................... 76

7.1.5. Launch Vehicle Processing Facilities ...................................................................................... 76

7.2. Payload Proces s in g/Int egr a tion .................................................................................................. 77

7.2.1. Payload Propellant Loading ..................................................................................................... 78

7.3. Launch Operations ...................................................................................................................... 78

7.3.1. Booster Assembly Stacking/Launch Site Preparation ............................................................. 78

7.3.2. Final Vehicle Integration and Test ........................................................................................... 79

7.3.3. Launch Vehicle Arming ............................................................................................................ 79

7.3.4. Launch ..................................................................................................................................... 80

7.3.5. Launch Control Organization ................................................................................................... 80

7.3.6. Launch Rehearsals .................................................................................................................. 81

8. OPTIONAL ENHANCED CAPABILITIES .......................................................................................... 82

8.1. Separation System and Optional Mechanical Interfaces ............................................................ 82

8.2. Conditione d Air ........................................................................................................................... 82

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8.3. Nitrogen Purg e ............................................................................................................................ 82

8.4. Additional Access Panel ............................................................................................................. 82

8.5. Enhanced Telemetry ................................................................................................................... 83

8.6. Enhanced Contamination Control ............................................................................................... 83

8.6.1. Low Outgassing Materials ....................................................................................................... 83

8.6.2. High Cleanliness Integration Environment .............................................................................. 84

8.6.3. HEPA-Filtered Fairing Air Supply ............................................................................................ 84

8.6.4. Fairing Surface Cleanliness ..................................................................................................... 84

8.7. Secure FTS ................................................................................................................................. 84

8.8. Over Horizon Telemetry .............................................................................................................. 85

8.9. Increased Insertion Accuracy ...................................................................................................... 86

8.10. Payload Isolation System ............................................................................................................ 87

8.11. Orbital Debris Mitigation .............................................................................................................. 87

8.12. Dual and Multi Payload Adapter Fittings ..................................................................................... 88

8.13. Enhanced Performance .............................................................................................................. 88

8.14. Large Fairing ............................................................................................................................... 88

8.15. Hydrazine Servici ng .................................................................................................................... 88

8.16. Nitrogen Tetroxide Service ......................................................................................................... 89

8.17. Poly-Pico Orbital Deployer (P-POD) ........................................................................................... 90

8.18. Suborbital Performance .............................................................................................................. 91

8.19. Alternate Launch Locations ........................................................................................................ 92

LIST OF FIGURES

Figure 1.1-1. The Minotaur Family of Launch Vehicles ............................................................................... 2

Figure 1.1-2. Space Launch Performance for the Minotaur Family Demonstrates a Wide Range of

Payload Lift Capability ............................................................................................................ 3

Figure 2.1-1. Minotaur IV Baseline Launch Vehicle ..................................................................................... 4

Figure 2.3-1. Minotaur IV Expanded View Showing Orbital’s State-of-the-Art

Structures and Modular Architecture ...................................................................................... 6

Figure 2.3.1-1. GFE Peacekeeper Stages 1, 2, and 3 Have an Extensive Flight History with

over 50 Launches ................................................................................................................ 6

Figure 2.3.2-1. Orion 38 Stage 4 Motor ....................................................................................................... 7

Figure 2.3.3-1. The Adaptable and Flexible Design of Minotaur Affords a Wide Range of Options for

Payload Customers .............................................................................................................. 7

Figure 2.3.5-1. Minotaur IV 92” Fairing and Handling Fixtures .................................................................. 10

Figure 2.3.6.1-1. Minotaur IV+ Enhancement ............................................................................................ 11

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Figure 2.3.6.2-1. Minotaur V Vehicle Configuration ................................................................................... 12

Figure 2.3.6.3-1. Minotaur VI Vehicle Configuration .................................................................................. 13

Figure 2.3.6.4-1. Minotaur VI+ Vehicle Configuration ................................................................................ 14

Figure 2.4-1. Multiple Sets of MGSE Are Availab le to Sup port Par a llel Miss io ns ..................................... 15

Figure 2.4-2. Functional Block Diagram of LSE ......................................................................................... 16

Figure 3.1-1. Generic Minotaur IV Mission Profile ..................................................................................... 17

Figure 3.2-1. Flexible Processing and Portable GSE Allows Operations from Multiple Ranges or

Austere Site Options ............................................................................................................. 18

Figure 3.2-2. Launch Site Inclinations ........................................................................................................ 18

Figure 3.3-1. Minotaur IV Fleet Comparison Performance Curves for SSO Out of KLC ........................... 21

Figure 3.3-2. Minotaur IV Fleet Comparison Performance Curves for 28.5° Inclination

Orbits Out of CCAFS ............................................................................................................ 21

Figure 3.3.1-1. Minotaur IV Performance Curves for VAFB Launches ...................................................... 22

Figure 3.3.1-2. Minotaur IV Performance Curves for KLC Launches ........................................................ 22

Figure 3.3.1-3. Minotaur IV Performance Curves for CCAFS Launches ................................................... 23

Figure 3.3.1-4. Minotaur IV Performance Curves for WFF Launches ....................................................... 23

Figure 3.3.2-1. Minotaur IV+ Performance Curves for VAFB Launches .................................................... 24

Figure 3.3.2-2. Minotaur IV+ Performance Curves for KLC Launches ...................................................... 24

Figure 3.3.2-3. Minotaur IV+ Performance Curves for CCAFS Launches ................................................. 25

Figure 3.3.2-4. Minotaur IV+ Performance Curves for WFF Launches ..................................................... 25

Figure 3.3.3-1. Minotaur VI (92” Fairing) Performance Curves for CCAFS Launches .............................. 26

Figure 3.3.3-2. Minotaur VI (92” Fairing) Performance Curves for KLC Launches ................................... 26

Figure 3.3.3-3. Minotaur VI (110” Fairing) Performance Curves for CCAFS Launches ............................ 27

Figure 3.3.3-4. Minotaur VI (110” Fairing) Performance Curves for KLC Launches ................................. 27

Figure 3.3.4-1. Minotaur IV+ Elliptical Orbits Performance Curve for KLC Launches ............................... 28

Figure 3.3.4-2. Minotaur V Elliptical Orbits Performance Curve for KLC Launches .................................. 28

Figure 3.3.4-3. Minotaur VI+ Elliptical Orbits Performance Curves for KLC Launches ............................. 29

Figure 3.3.4-4. Minotaur V/VI+ High Energy Orbit Performance Curves for CCAFS Launches................ 29

Figure 3.3.4-5. Minotaur V High Energy Orbit Performance Curve for WFF Launches ............................ 30

Figure 4.1.1-1. Payload CG Net Transient Lateral Acceleration (Minotaur IV) .......................................... 35

Figure 4.1.1-2. Payload CG Net Transient Lateral Acceleration (Minotaur IV+, V, VI, and VI+) ............... 35

Figure 4.1.2-1. Minotaur IV Family Maximum Axial Acceleration as a Function of Payload Mass ............ 36

Figure 4.2.1-1. Minotaur IV Family Payload Random Vibration Environment ........................................... 37

Figure 4.2.2-1. Minotaur IV, IV+, and VI Payload Sine Vibr ati on MP E Lev els .......................................... 38

Figure 4.2.2-2. Minotaur V and VI+ Paylo ad Sine Vi bration MPE Levels .................................................. 38

Figure 4.3-1. Minotaur IV Payload Acoustic Maximum Predicted Environment (MPE) with

1/3 Octave Breakpoints ......................................................................................................... 39

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Figure 4.4-1. Minotaur IV Family Payload Shock Maximum Predicted Environment (MPE) –

Launch Vehicle to Payload ................................................................................................... 40

Figure 4.4-2. Minotaur V/VI+ Payload Shock MPE – Launch Vehicl e to Pa yload ..................................... 41

Figure 4.4-3. Maximum Shock Environment - P a ylo ad to Launch Ve hic l e ................................................ 41

Figure 4.6.1-1. Minotaur IV HVAC System Provides Conditioned Air to the Payload ............................... 43

Figure 5.1.1.1-1. Dynamic Envelope for Standard 92” Fairing with Standard 38” PAF ............................. 46

Figure 5.1.2.1-1. Dynamic Envelope for Optional 110” Fairing with Standard 38” PAF ............................ 47

Figure 5.1.3-1. Available Fairing Access Door Locations .......................................................................... 48

Figure 5.2.1-1. Minotaur IV Coordinate System......................................................................................... 49

Figure 5.2.3-1. Standard, Non-separating 38.81” Diameter Payload Mechanical Interface ...................... 50

Figure 5.2.4.1-1. Dynamic Envelope for Sta ndard 92” Fair in g with

Optional 38” 2-Piece Payload Cone ............................................................................... 53

Figure 5.2.4.1-2. Dynamic Envelope for Standard 92” Fairing with Optional 62” Payload Cone ............. 53

Figure 5.2.4.1-3. Dynamic Envelope for Optional 110” Fairing with Optional 62” Payload Cone .............. 54

Figure 5.2.4.2.1-1. Dynamic Envelope for Standard 92” Fairing with Optional DPAF .............................. 54

Figure 5.2.4.2.2-1. Dynamic Envelope for Standard 92” Fairing with Optional MPAF .............................. 55

Figure 5.2.4.2.2-2. Optional Multi-Payload Adapter Plate (MPAP) Non-Separating Mechanic a l

Interface – Accommodates 2 to 4 ESPA-Class Payloads ........................................... 55

Figure 5.2.4.2.3-1. Dynamic Envelope for Standard 92” Fairing and Minotaur V / VI+

Enhanced Performance Option ................................................................................... 56

Figure 5.2.4.2.3-2. Dynamic Envelope for Standard 92” Fairing and Minotaur V / VI+

Enhanced Performance Option with Optional PAF .................................................... 56

Figure 5.2.4.2.3-3. Minotaur V PAF Non-Separating Mechanical Interface .............................................. 57

Figure 5.2.5.1-1. Orbital 38 ” Separati on S yst em ....................................................................................... 58

Figure 5.2.5.2-1. 38” Planetary Sciences Motorized Lightband ................................................................. 59

Figure 5.2.5.3-1. RUAG 937S 38” Separation System .............................................................................. 59

Figure 5.3.1-1. Payload 1:1 Umbilical Pin Outs ......................................................................................... 60

Figure 5.3.2-1. Payload Electrical Interface Block Diagram ...................................................................... 60

Figure 6.1-1. Mission Integration Team ..................................................................................................... 64

Figure 6.2-1. Typical Mission Integration Schedule ................................................................................... 66

Figure 6.2-2. Typical Mission Field Integration Schedule .......................................................................... 67

Figure 7-1. Hardware Flow – Factory to Launch Site ................................................................................ 74

Figure 7.1-1. Launch Vehicle Processing Flow at the MPF ....................................................................... 75

Figure 7.1-2. Minotaur Launch Vehicle Integration at MPF ....................................................................... 75

Figure 7.1.5-1. Minotaur Processing Is Performed at the MPF at VAFB ................................................... 76

Figure 7.2-1. Payload Processing and LV Integration Flow at the PPF .................................................... 77

Figure 7.2-2. Payload Encapsulation at the PPF ....................................................................................... 77

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Figure 7.2-3. Encapsulated Payload Transport to the Launch Site ........................................................... 77

Figure 7.3-1. Minotaur IV Launch Site Operations .................................................................................... 78

Figure 7.3.1-1. Minotaur Uses Vertical Integration for Each Booster Stage, the Guidance Control

Assembly, and the Encapsulated Pa yload As sembly ....................................................... 79

Figure 7.3.4-1. Notional Minotaur Countdown Timeline ............................................................................ 80

Figure 7.3.4-2. Minotaur IV Prepared for Launch ...................................................................................... 80

Figure 8.3-1. GN2 Purge Interface To Minotaur Fairing (Flyaway at Liftoff) .............................................. 82

Figure 8.4-1. Example Location and Size of Additional Access Panel ...................................................... 82

Figure 8.5-1. Representative Minotaur 92” Enhanced Instrumentation Locations (Fairing Only) ............. 83

Figure 8.6-1. Minotaur Team Has Extensive Experience in a Payload Processing

Clean Room Environment ..................................................................................................... 84

Figure 8.7-1. Orbital’s Secure FTS System Block Diagram ....................................................................... 85

Figure 8.8-1. TDRSS 20W LCT2 Transmitter and Cavity Backed S-band Antenna .................................. 85

Figure 8.8-2. TDRSS Notional Telemetry Flow.......................................................................................... 86

Figure 8.9-1. Hydrazine Auxiliary Propulsion System (HAPS) Used to Provide Insertion Accuracy ......... 86

Figure 8.10-1. Minotaur Soft Ride Significantly Attenuates Peak LV Dynamic Environments .................. 87

Figure 8.11-1. Operational and Disposal LEOs ......................................................................................... 87

Figure 8.15-1. Typical Propellant Loading Schematic ............................................................................... 88

Figure 8.15-2. UPC Provides Reliable and Demonstrated Hydrazine Servicing for Minotaur ................... 89

Figure 8.17-1. P-PODs Have Successfully Flown On Multiple Minotaur Missions .................................... 90

Figure 8.18-1. Expanded View of Minotaur IV Lite Configuration .............................................................. 91

Figure 8.19-1. Minotaur IV Vehicle Processing and Launch From KLC .................................................... 92

Figure 8.19-2. Launch Complex 46 at CCAFS Supports All Minotaur Configurations .............................. 92

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LIST OF TABLES

Table 3.2-1. Baseline Launch Sites for the Minotaur IV Family of Launch Vehicles ................................. 17

Table 3.3-1. Common Mission Options and Associated Masses (These Masses Must Be

Subtracted from the LV Performance) ................................................................................... 20

Table 3.3.4-1. Geosynchronous Transfer Orbit (GTO) Performance For CCAFS ..................................... 31

Table 3.3.4-2. Medium Transfer Orbit (MTO) Performance For CCAFS ................................................... 31

Table 3.3.4-3. Medium Transfer Orbit (MTO) Performance For WFF ....................................................... 31

Table 3.4-1. Minotaur IV Injection Accuracy .............................................................................................. 32

Table 3.5-1. Typical Pre-Separation Payload Pointing and Spin Rate Accuracies .................................. 32

Table 4.8-1. Minotaur IV Launch Vehicle RF Emitters and Receivers ...................................................... 45

Table 5.2.4-1. Minotaur IV Payload Adapter Fitting Options ..................................................................... 51

Table 5.2.5-1. Minotaur IV Separation System Options ............................................................................ 57

Table 5.4.2-1. Payload Mass Properties Measurement Tolerance ........................................................... 62

Table 8.9-1. Enhanced Insertion Accuracy ................................................................................................ 86

LIST OF APPENDICES

A. PAYLOAD QUESTIONNAIRE ..............................................................................................................A-1

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Minotaur IV • V • VI User’s Guide Glossary

A-CAT 1

Acquisition Category

A/D

Arm/Disarm

AADC

Alaska Aerospace Development
ACS

Attitude Control System

AFRL

Air Force Research Laboratory

ait

Atmospheric Interceptor Technology

BCM

Booster Control Module

BER

Bit Error Rate

C/CAM

Collision/Contamination Avoidance
CBOB

Clamp Band Opening Device

CCAFS

Cape Canaveral Air Force Station

CDR

Critical Design Review

CG

Center-of-Gravity

CLA

Coupled Loads Analysis

CLF

Commercial Launch Facility

CVCM

Collected Volatile Condensable

DIACAP

DoD Information Assurance

DPAF

Dual Payload Adapter Fitting

ECU

Environmental Control Unit

EELV

Evolved Expendable Launch Vehicle

EGSE

Electrical Ground Support
EME

Electromagnetic Environment

EMI

Electromagnetic Interface

ER

Eastern Range

ESPA

EELV Secondary Payload Adapter

FRR

Flight Readiness Review

FTS

Flight Termination System

GCA

Guidance and Control Assembly

GEO

Geostationary Earth Orbit

GFE

Government Furnished Equipment

GFP

Government Furnished Property

GN2

Gaseous Nitrogen

GPB

GPS Positioning Beacon

GPS

Global Positioning System

GTO

Geosynchronous Transfer Orbit

HAPS

Hydrazine Auxiliary Propulsion

HVAC

Heating, Ventilation and Air
ICD

Interface Control Document

INS

Inertial Navigation System

KLC

Kodiak Launch Complex

KSC

Kennedy Space Center

LAN

Longitude of Ascending Node

LC-46

Launch Complex 46

LCR

Launch Control Room

LEO

Low Earth Orbit

LEV

Launch Equipment Vault

LOCC

Launch Operations Control Center

LRR

Launch Readiness Review

LSE

Launch Support Equipment

LSV

Launch Support Van

LV

Launch Vehicle

MACH

Modular Avionics Control Hardware

MARS

Mid-Atlantic Regional Spaceport

MDR

Mission Design Review

MDR

Mission Dress Rehearsal

MGSE

Mechanical Ground Support

MICD

Mechanical Interface Control
MIWGs

Mission Integration Working Groups

MLB

Motorized Lightband

MPAF

Multi-Payload Adapter Fitting

MPAP

Multi-Payload Adapter Plat e

MPE

Maximum Predicted Environment

MPF

Minotaur Processing Facility

MRD

Mission Requirements Document

MRR

Mission Readiness Review

MST

Mission Simulation Test

MTO

Medium Transfer Orbit

NTO

Nitrogen Tetroxide

ODM

Ordnance Driver Module

OML

Outer Mold Line

OR

Operations Requirements

OSP-3

Orbital Suborbital Program 3

P-POD

Poly-Pico Orbital Deployer

PAF

Payload Attach Fitting

PAM

Payload Adapter Module

PCM

Pulse Code Modulation

PDR

Preliminary Design Review

PEM

Program Engineering Manager

PPF

Payload Processing Facility

PRD

Program Requirements Document

RAAN

Right Ascension of Ascending Node

RF

Radio Frequency

Corporation

Maneuver

Mass

Certification and Accreditati on
Process

Equipment

Equipment

Drawing

System

Conditioning

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RTS

Range Tracking System

RWG

Range Working Group

SD

Space Development and Test
SDL

SD Launch System Division

SEB

Support Equipment Building

SLC-8

Space Launch Complex 8

SCAPE

Self-Contained Atm os pheri c
SEB

Support Equipment Building

SRSS

Softride for Small Satellites

SSI

Spaceport Systems International

START

Strategic Arms Reduction Treaty

TDRSS

Telemetry Data Relay Satellite
TML

Total Mass Loss

TVA

Thrust Vector Actuator

TVC

Thrust Vector Control

UPC

United Paradyne Corporation

VAFB

Vandenberg Air Force Base

WFF

Wallops Flight Facility

WPs

Work Packages

Directorate

Protective Ensemble

System

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Minotaur IV • V • VI User’s Guide Section 1.0 - Introduction

1. INTRODUCTION

This User’s Guide is intended to familiarize payload
mission planners with the capabilities of the Orbital
Suborbital Program 3 (OSP-3) Minotaur IV Launch
Vehicle (LV) launch ser vice. The inf ormation provided
in this user’s guide is for initial planning purposes
only. Information for development/design is
determined through mission specific engineering
analyses. The results of these analyses are
documented in a mission-specific Interface Control
Document (ICD) f or the p ayload organizati on t o us e in
their development/des ign process. This User’s Guide
provides an overview of the Minotaur IV family of
launch vehicles system design and a description of
the services provided to ou r customers. T he Minotaur
IV family of launch vehic les includes the Minot aur IV,
IV+, V, VI, and VI+. Minotaur vehicles off er a variety
of enhancement options to allow the maximum
flexibility in satisfying the objectives of single or
multiple payloads.

The primary mission of the Minotaur IV family of
vehicles is to provide low cost, high rel iability launch
services to government-sponsored payloads. The
Minotaur design accomplishes this using flight prov en
components with significant flight heritage. The
philosophy of placin g mission success as the highest
priority is reflected in the success and accuracy of all
Minotaur missions to date.

This User’s Guide des cr ib e s the b as ic e lements of the
Minotaur IV system as well as optional services that
are available. In addition, this document provides
general vehicle performance, defines payload
accommodations and envir onments, and outlines the
Minotaur mission integration process. Minotaur­unique integration and t est approaches (includi ng the
typical operational timeline for payload integration
with the Minotaur vehicles) and the ground support
equipment that will be used to conduct Minotaur
operations are also described.

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1.1. Minotaur Family Performance and Capability

Figure 1.1-1 shows the Min otaur fam ily of la unch vehic les, w hich is capable of lau nching a wide rang e of
payload sizes and missions. Representative space launch performance across the Minotaur fleet is
shown in Figure 1.1-2 and illustrates the relative capability of each c onfiguration. In addition to space
launch capabilities, t he M in otaur I Lit e an d M ino taur I V Lite co nf igur at io ns ar e a va i lab le to meet suborbital
payload needs for payloads weighing up to 3000 kg. T his User’s Guide covers t he Peacekeeper-based
Minotaur IV family. Please refer to the Minotaur I User’s Guide for information regarding Minuteman­based Minotaur vehicles.

Figure 1.1-1. The Minotaur Family of Launch Vehicles

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Figure 1.1-2. Space Launch Performance for the Minotaur Family Demonstrates a Wide Range of

Payload Lift Capability

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2. MINOTAUR IV CONFIGURATIONS

2.1. Minotaur IV Launch System Overview

The Minotaur IV (Figure 2.1-1) mission is to
provide a cost effective, reliable and flexible
means of placing s atellites into orbit. O rbita l is the
launch vehicle provider and manufacturer under
the Orbital Suborbital Program 3 (OSP-3) contract
for the U.S. Air Force. An overview of the system
and available launch services is provided within
this section, with specific elements covered in
greater detail in the subsequent sections of this
User’s Guide.

The Minotaur IV family of launch vehicles has
been designed to meet the needs of U.S.
Government-sponsore d cus tomers at a lower cos t
than commercially av ailable alternatives by using
surplus Peacekeeper boosters. As stated
previously, the Minotaur IV family of launch
vehicles includes the Minotaur IV, IV+, V, VI, and
VI+. The requirements of the OSP-3 program
emphasize system reliability, transportability, and
operation from multiple launch sites. Minotaur IV
draws on the successful heritage of Orbital’s
space launch vehicles as well as the USAF
Peacekeeper system to m eet these requir ements.
Orbital has built upon thes e legacy systems with
enhanced avionics components and advanced
composite s tructures to meet the payload-support
requirements of the OSP-3 program. Combining
these subsystems with the long successf ul history
of the Peacekeeper boosters has resulted in a
simple, robust, self-contained launch system to
support government-sponsored small satellite
launches.

The Minotaur IV system also includes a com plete
set of transportable Launch Support Equipment
(LSE) designed to allow Minotaur IV to be
operated as a self-contained satellite delivery
system. To accomplish this goal, the Electrical
Ground Support Equipment (EGSE) has been
developed to be portab le a nd a dapt able to varying
levels of infrastructure. While the Minotaur IV
system is capable of self-contained operation at

Figure 2.1-1. Minotaur IV Baseline

Launch Vehicle

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austere launch sites using portable vans, typical operations occur from permanent facilities on
established ranges.

The Minotaur IV s ystem is design ed to be capable of launch from four commercial Spaceports (Alask a,
California, Florida, and Mid-Atlantic), as well as from existing U.S. Government facilities at VAFB and
CCAFS. A Launch Control Room (LCR) serves as the control center f or conducting a Minotaur IV launch
and includes consoles f or Orbital , r ang e s af et y, and limited custom er pers onnel. Fur ther des cr ipt ion of the
Launch Support Equipment is provided in Section 2.4.

2.2. Minotaur IV Launch Service

The Minotaur IV L aunch S ervice is pr ovided throu gh the com bined eff orts of the USAF a nd Orbital, alo ng
with associate contractors and commercial spaceport s. The primary customer interface will be with the
USAF Space and Missile Systems Center, Space Development and Test Directorate (SD), Launch
Systems Division (SDL). Orbital is the la unch vehicle provider. This integra ted team will be referred to
collectively as “OSP” throughou t the User’s Gu ide. Where necessary, interfaces that are associat ed with
a particular member of the team will be referred to directly (i.e., Orbital or SDL).

OSP provides all of the necessary hardware, software and services to integrate, test, and launch a
payload int o its prescribed orbit. In addition, O SP will complete al l the re quired agr eements, licenses and
documentation to successfully conduct Minotaur IV operations. The Minotaur IV mission integration
process completely identifies, documents, and verifies all spacecraft and mission requirements.

2.3. Minotaur IV Launch Vehicle

The Minotaur IV baseline vehicle, shown in expanded view in Figure 2.3-1, is a four-stage, inertially
guided, all solid propellant ground launched vehicle. Conservative design margins, state-of-the-art
structural systems , a modular avionics architecture and s implified integration and test cap ability yield a
robust, highly reliable launch vehicle design. In addition, Minotaur IV payload accommodations and
interfaces are designed to satisfy a wide range of potential payload requirements.

2.3.1. Stage 1, 2 and 3 Booster Assemblies

The first three stages of the Minotaur IV consist of the refurbished Government Furnished Equipment
(GFE) Peacekeeper Sta ges 1, 2, and 3, sho wn in Figure 2.3.1-1. Thes e booster ass emblies are used as
provided by the Governm ent, requiring no modification. They have extensive flight his tory, with over 50
launches. All three sta ges are s olid-propellant motors and utili ze a movable nozzle c ontrolled b y a Thrus t
Vector Actuator (TVA) s ystem for three-axis attitude control. The first stage provides 500 ,000 lbf (2224
kN) of thrust. T he second stage motor has an ext endable exit cone and provides an avera ge thrust of
275,000 lbf (1223 kN). The third stage provides 65,000 lbf (289 kN) of thrust and also features an
extendable exit cone similar to Stage 2.

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Flight History with over 50 Launches

Figure 2.3-1. Minotaur IV Expanded View Showing Orbital’s State-of-the-Art

Structures and Modular Architecture

2.3.2. Upper Stage Propulsion

The Minotaur IV baseline Stage 4 motor is the Orion 38 (Figures 2.3.2-1). This motor was originally
developed for Orbit al’s Pegasus program and is used on many other Orbita l launch vehicles, inc luding
Minotaur I. The Orion 38 motor provides the velocit y needed for orbit insert ion for the launch vehicle, in
the same manner as it is u s ed on t he M ino taur I. This motor features state-of-the-art design and m aterials
with a successful flight heri t age. I t is c urr ent l y in production and is activel y fl ying p a yloa ds into spac e , with
over 60 launches.

th

While the baseline Minotaur IV 4

Stage is the
Orion 38, the flexible Minotaur IV design allows for
a number of performance enhanc ements such as
replacing the Orion 38 with a STAR 48, adding a

th

stage ST AR 37 m otor , and a dding a Hydrazine

5
Auxiliary Propulsion System (HAPS) for precise
targeting or orbital insertion requirements. These
options are described i n detail later in this section
as well as in Section 8.0.

Figure 2.3.1-1. GFE Peacekeeper

Stages 1, 2, and 3 Have an Extensive

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a Wide Range of Options for Payload Customers

2.3.3. Guidance and Control Assembly (GCA)

The Guidance and C ontrol As sembl y (GCA) is the
heart of the launch vehicle, comprised of an
Avionics Assembl y as well as the GCA Skirt which
forms the 92 inch Outer Mold Line (OML). The
Avionics Assembly houses all of the required
subsystems for vehicle operation including power,
telemetry, RF, ordnance, pneumatic, and
guidance and control. In a ddit io n, th e an nu lar ring
design of the Avionics Ass embly enables multiple
upper stage motor options (Figure 2.3.3-1). The
GCA skirt has four large doors for ease of access
to components within the Avionics Assembly.
Antennas and thruster ports are mounted to the
GCA skirt to allow for clear and unimpeded
operation during flight.

Figure 2.3.2-1. Orion 38 Stage 4 Motor

Figure 2.3.3-1. The Adaptable and Flexible Design of Minotaur Affords

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2.3.3.1. Avionics

The Minotaur avionics system has heritage and commonality across the Minotaur fleet. The flight
computer is a 32-bit multipr ocessor arc hitecture. It pro vides comm unication w ith vehicle subs ystems , the
LSE, and if required, the payload via standard RS-422 serial links and discrete I/O. T he avionics s ystem
design incorporates Orbital’s innovative, flight proven Modu lar Avionics Control Hardware (M ACH). The
MACH consists of standardized, function-specific modules that are combined in stacks of up to 10
modules to m eet mission requirem ents . T he functional m odules from which the M ACH stac k s are created
include power trans fer, ordnance initiation, booster interface, com munication, and telem etry processing.
These modules provide an array of functional capability and flexibility in mission tailoring.

2.3.3.2. Attitude Control Systems

The Minotaur IV Control System provides attitude control throughout both boosted flight and coast
phases. The Orbital-developed Booster Control Module (BCM) links the flight computer actuator
commands to the individual Thrust Vector Actuators (TVAs) located on each PK motor. The available
upper stage motors (Orion 38, STAR 48 and STAR 37) are commanded with the sam e Thrust Vector
Control (TVC) control methodology as Minotaur I. This control combines a single-nozzle
electromechanical T VC for pitch and yaw augmented with roll control from a three-axis, cold-gas Attitude
Control System (ACS) resident within the GCA. The cold-gas ACS also provides 3-axis control as
necessary during exoatmospheric coast and post-boost phases of flight.

Attitude control is achieved us ing a three-axis autopilot. Stag es 1, 2 an d 3 fly a p re-programm ed attitude
profile based on trajectory design and optimization. Stage 4 uses a set of pre-programmed orbital
parameters to place the vehicl e on a trajectory toward the intended insertion a pse. An extended coast
between Stages 3 an d 4 is used to or ient the vehicle t o the appropr iate att itude f or Stage 4 ign ition bas ed
upon a set of pre-programmed orbital parameters and the measured performance of the first three
stages. Stage 4 utilizes ene r g y manag ement to place the vehicle into the pr o per o r bit. Af ter t he f ina l boost
phase, the three-axis cold-gas attitude control system is used to orient the vehicle for spacecraft
separation, contamination and collision avoidance and downrange downlink maneuvers. The autopilot
design is modular, so a dditional payload requirement s such as rate control or celestial point ing can be
accommodated with minimal development effort.

2.3.3.3. Telemetry Subsystem

The Minotaur IV telemetry subs ystem provides real-t ime health and status data of the vehic le avionics
system, as well as k ey inform ation regarding the position, perf ormance and envir onment of the Minotaur
IV vehicle. This data is used by both Orbital and the range safety personnel to evaluate system
performance. The Minotaur IV baseline telemetry subsystem provides a number of dedicated payload
discrete (bi-level) and analog telemetry monitors through dedicated channels in the launch vehicle
encoder. The baseline te lemetry system has a 1.5 Mbps data rate for both pa yload and Minotaur launch
vehicle telemetry. To allow for flexibilit y in support ing evolv ing m ission requirem ents, the output data rat e
can be selected over a wide range from 2.5 kbps to 10 Mbps (c ontingent on link margin and Bit Error
Rate (BER) requirements). The telemetry subsystem nominally utilizes Pulse Code Modulation (PCM)
with a RNRZ-L format. Other types of data formats, including NRZ-L, S, M, and Bi-phase may be
implemented if required to accommodate launch range limitations. Furthermore, the launch vehicle
telemetry system has the capability to take payload telemetry as an input, randomize if required, and
downlink that dedicated payload link from launch through separation. That capability is available as a
non-standard option.

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Module

The Enhanced Telemetry option as described in the Enhancements section 8.5 augments the existing
baseline telemetr y system by provid ing a dedicated telem etry link with a baseline data rate of 2 Mbps.
This Enhanced Telemetry link is used to provide further insight into the mission environment due to
additional payload, LV, or experiment data acquisition requirements. Supplementary instrumentation or
signals such as strain gauges, temperature sensors, accelerometers, analog or digital data can be
configured to meet payload mission-specific requirements.

An Over The Horizon T elem etry option can also be added to provid e real-tim e telemetr y coverag e during
ground-based telemetry receiving site blackout periods. The Telemetry Data Relay Satellite System
(TDRSS) is use d for this capab ility, and h as been s uccessfull y demonstrated on past Minot aur miss ions.
Close to the time when telemetry coverage is lost by ground based telemetry receiving sites, the LV
switches telemetry output to the TDRSS ante nna and points the antenn a towards the designat ed satellite.
The TDRSS then rela ys the telemetr y to the ground where it is r outed to the laun ch control room for real­time telemetry updates . Reference Enhancem ents section 8.8 f or f urther detai ls o n this Over T he Hori zon
Telemetry option.

Minotaur telemetr y is subject to the provisions of the Strategic Arms Reduction Treaty (START ). START
treaty provisions require that certain Minotaur telemetry be unencrypted and provided to the START
treaty office for dissemination to the signatories of the treaty.

2.3.4. Payload Interface

Forward of the GCA is the Payload Adapter Module (P AM) , shown in Figure 2.3.4-1. It is c om pr ised of the
fairing frangible separation ring, fairing adapter ring and payload cone, which adapts from the 92 inch
OML down to the standard 38 inch i nterface. This assem bly provides both the mec hanical interface with
the payload as well as serves to close out the bottom of the encapsulated payload volume.

Minotaur provides f or a standard non-separating payload interf ace with the option of add ing an Orbital­provided payload separation system. Orbital will provide all flight hardware and integration services
necessary to attach non-separ ating and separating payloads to the Minotaur launch vehicle. Additional
mechanical interface diameters and separation system configurations can readily be provided as an
enhanced option as described in Section 5.0. Further detail on p ayload e lectrical, mec hanical and launc h
support equipment interfaces can also be found in Section 5.0.

With the addition of various structural adapters,
the Minotaur IV can accommodate multiple
payloads. This featur e, des cribed in m ore detail in
Section 5.2.4.2 of this User’s Guide, permits two
or more payloads to share the c ost of a Minotaur
IV launch, thus lowering the launch cost when
compared to other launch options. Furthermore,
Orbital can accommodate small payloads when
there is excess payload and/or mass capability.

Figure 2.3.4-1. Minotaur IV Payload Adapter

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2.3.5. Payload Fairing

Orbital’s flight proven Minotaur IV 92” diameter
payload fairing (Figure 2.3.5-1) is used to
encapsulate the payload, providing protection and
contamination control during ground handling,
integration operations and flight. The fairing is a
bi-conic design made of graphite/epoxy face
sheets with an aluminum honeycomb core. The
fairing provides for low payload contamination
through prudent design and selection of low
contamination materials and processes. Acoustic
blankets and internal c onditioned air are stan dard
service items that provide a m ore benign payload
environment. Conditioned air will keep the
payload environment to a specified temperature
between 13 to 29 °C (55 to 85 °F) dependent
upon requirements.

The two halves of the fairing ar e structurally joined along their longit udinal interface using Orbital’s low
contamination frang ible join t system . An additional cir cumf erential frangi ble joint at the base of the fair ing
supports the fairing loads. At separation, a cold-gas system is activated to pressurize the fairing
deployment thrusters. The fairing halves then rotate about external hinges that control the fairing
deployment to ensure that payload and launch vehicle clearances are maintained. All elements of the
deployment system have been demonstrated through numerous ground tests and flights.

The Minotaur IV comes standard with a single pa yload access door; however, options for extra payload
access doors and enhanc ed cle anlin ess ar e avai lable . Fur ther detai ls on th e bas eli ne f airing are inc luded
in Section 5.1.

A larger 110” diameter fairing design is available as an enhancement (reference Section 5.1.2) to
accommodate payloads l arger than those t hat can be f it in the standard 92” d iameter Minotaur IV f airing.
The fairing, composite materials, structural testing, separation and deployment systems are similar to
those of the heritage 92” fairing.

Figure 2.3.5-1. Minotaur IV 92” Fairing and

Handling Fixtures

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2.3.6. Minotaur IV Launch Vehicle Enhanced Performance Configurations

The modular design of the Minotaur IV vehicle
allows for a substantial increase in performance
with minimal vehicle changes. The Minotaur IV
Enhanced Perform ance Configurations ut ilize the
identical flight proven Peacekeeper stages,
mechanical structures, avionics hardware,
mechanical pneumatics, and ordnance
subsystems as the base Minotaur IV vehicle.

The Minotaur IV Enhanced Performance
Configurations are built to the same stringent
requirements as the Minotaur IV vehicle and
undergo an identical rigorous testing program.

2.3.6.1. Minotaur IV+ (STAR 48 Stage 4)

The flight proven Minotaur IV+ vehicle, shown in
Figure 2.3.6.1-1, utilizes the larger STAR 48BV
motor in place of the standard Stage 4 Orion 38
motor. Minotaur IV+ provides approximately 200
kg of increased perf ormance to low-earth circular
orbits and enables missions requiring highly
elliptical orbits. T he M in ota ur IV + vehic l e is ab le to
offer this increased performance without
sacrificing available payload volume.

The adjustments necessary for the Minotaur IV+
only require the exchange of the standard Orion
38 composite Motor Adapter Cone for the STAR
48BV Motor Adapter Cone and the addition of a
short extension structure to allow for the
increased motor length.

The STAR 48BV provides an average burn time of

85.2 seconds at an average thrust of 68.63 kN
(15.43 k-lbf). The total STAR 48BV mass is
2171 kg (4777 lbm), including a propellant mass
of 2014 kg (4431 lbm).

Figure 2.3.6.1-1. Minotaur IV+ Enhancement

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2.3.6.2. Minotaur V (High Energy Performance)

The Minotaur V vehicle is a five stage evo lutionar y
version of the Minotaur IV vehicl e, sho wn in Figure

2.3.6.2-1, which provides a cost-effective
capability to place small spacecraft into high
energy trajectories, including Geosynchronous
Transfer Orbit (GTO), Medium Transfer Orbit
(MTO), as well as translunar injection. The
Minotaur V vehicle leverages Orbital’s flight proven
heritage of the Minotaur IV and IV+ vehicles.

Minotaur V builds upon the Minotaur IV+
enhancement by incorporating a STAR 37 fifth
stage within the fairing envelope. The design
accommodates either spin-stabilized or 3-axis
controlled versions of the STAR 37. The Minotaur
V configuration represents a more than 25%
increase in performance for highly elliptical orbits.

The STAR 37FM provides an average b urn tim e of

62.5 seconds at an average thrust of 48.13 kN
(10.82 k-lbf) and a total impulse of 3048 kN-sec
(685.4 lbf-sec). The total STAR 37FM mass is
1150 kg (2531 lbm), including a pr opell ant m ass of
1068 kg (2350 lbm).

Figure 2.3.6.2-1. Minotaur V Vehicle

Configuration

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2.3.6.3. Minotaur VI

The Minotaur VI launch vehicle, shown in Figure

2.3.6.3-1, provides 1800-3200 kg (4000-7000 lbm )
to Low Earth Orbit (LEO). Minotaur VI is a natura l,
low-risk evolution to the successful Minotaur IV
vehicle family, adding an additional Peacekeeper
Stage 1 (SR118) below the existing Minotaur IV+
stack (SR118-SR119-SR120-STAR 48BV). The
new design elements of Minotaur VI are based on
existing components, thereby minimizing risk.

Minotaur VI leverages heavily off the successful
Minotaur IV+ veh icle, using the same front sec tion
assembly. All avionic s, ordnance, an d pneumatics
components are already qualified for Minotaur VI
environments. All m ec hanic al s tr uctur es h av e be en
designed and qualified to loads that encompass
Minotaur VI with the exception of the payload
interface cone. However, payload cone
qualification is deemed low risk due to the safety
margins predicted for Minotaur VI loads. Minotaur
VI does not require new support equipment and
only requires minor procedural changes to use
existing Minotaur IV equipment and processes for
integration and test activities.

Existing facilities and infrastructure at Kodiak
Launch Complex (KLC) and Launch Complex 46
(LC-46) at CCAFS can accom modate Minotaur VI.
The Minotaur VI launch system complies with
range safety requirements RCC-319 and EWR
127-1, as tailored for Minotaur IV.

Figure 2.3.6.3-1. Minotaur VI Vehicle

Configuration

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2.3.6.4. Minotaur VI+ (High Energy Performance)

The Minotaur VI+ vehicle, shown in Figure

2.3.6.4-1, increases the Minotaur VI capability by
adding a STAR 37FM as the final stage. Minotaur
VI+ extends the Minotaur VI LEO capability to
3360 kg (7400 lbm). The Minotaur VI+ vehicle is
also very capable for highly elliptical orbits or Earth
escape trajectories such as a 300 kg (660 lbm)
spacecraft on a trajectory to Mars.

Similar to Minotaur V, Minotaur VI+ adds a STAR
37 stage within the fairing envelope. The design
accommodates either a spin-stabilized or 3-axis
controlled version of the STAR 37.

Existing facilities and infrastructure at Kodiak
Launch Complex (K LC) and LC-46 at CCAFS can
accommodate Minotaur VI. The Minotaur VI
launch system complies with range safety
requirements RCC-319 and EWR 127-1, as
tailored for Minotaur IV operations.

2.4. Launch Support Equipment

The Minotaur IV LSE is designed to be readily
adaptable to varying launch site configurations
with minimal unique infrastructure required. All of
the Mechanical Ground Support Equipment
(MGSE) used to support the Minotaur integr ation,
test, and launch is currently in use and launch
demonstrated, as shown in Figure 2.4-1. MGSE
fully supports all Minotaur configurations and are
routinely static load tested to safety factors in
compliance with Orbital internal and Range
requirements. The EGSE consists of readily
transportable consoles that can be housed in
various facility configurations depending on the
launch site infrastruc ture. The EGSE is composed
of three primary functional elements: Launch
Control, Vehicle Interface, and Telemetry Data
Reduction. The Launch Control and Telemetry
Data Reduction consoles are located in the
Launch Control Room (LCR), or mobile launch
equipment van dependi ng on available launc h site
accommodations. The Vehicle Interface consoles
are located at the launch pad in a permanent

Figure 2.3.6.4-1. Minotaur VI+ Vehicle

Configuration

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structure, typicall y called a Launch Equi pm ent Vault
(LEV). Fiber optic connections from the Launch
Control to the Vehicl e Interface consoles are used
for efficient, high bandwidth communications,
eliminating the need for copper wire between
locations. The Vehicle Interface consoles provide
the junction from fiber optic cables to the cables
that directly interf ace with the vehicle. Figure 2.4-2
depicts the functional bloc k diagram of the LSE. All
Minotaur EGSE is com pliant with th e D epartm ent of
Defense Instruction 8510.01, DoD Information
Assurance Certification and Accreditation Process
(DIACAP). Some launch sites have a separate
Support Equipment Building (SEB) that can
accommodate additional payload equipment.

The LCR serves as the control center during the
launch countdown. The number of personnel that
can be accommodated is dependent on the l aunch
site facilities. At a minimum, the LCR will accommodate Orbital personnel controlling the vehicle, two
Range Safety representatives (ground and flight safety), and the Air Force Mission Manager. Mission­unique customer-supplied payload consoles can be supported in the LCR, and payload equipment
required at the launch p ad can be suppor ted in the LE V or SEB, if available, within the constraint s of the
launch site facilities. Interface to the payload through the Minotaur IV payload umbilicals provides the
capability for direct monitor ing of payload functions. P ayload personnel accom modations will be handled
on a mission-specific basis.

Figure 2.4-1. Multiple Sets of MGSE Are

Available to Support Parallel Missions

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Figure 2.4-2. Functional Block Diagram of LSE

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Table 3.2-1. Baseline Launch Sites for the

Launch Vehicle

Baseline Launch Site

Minotaur IV/IV+

VAFB

Minotaur V

WFF

Minotaur VI/VI+

KLC

3. GENERAL PERFORMANCE

3.1. Mission Profiles

Minotaur IV fam ily of Launc h Vehicles can attain a range of pos igrade and r etro grade inc linations through
the choice of launch sites m ade available b y the readily adaptable n ature of the Minotaur launch system.
A generic mission profile to a sun-synchronous orbit is s ho wn in Figure 3.1-1. All performance parameters
presented within this User ’s Guide are typical for most expected payloads. However, perf ormance may
vary depending on unique payload or mission characteristics. Specific requirements for a particular
mission should be coord inated with OSP. Once a mission is f ormally initiated, the requirements will be
documented in the Mission Requirements Document (MRD). Further detail will be captured in the
Payload-to-Launch Vehicle Interface Control Document (ICD).

3.2. Launch Sites

Depending on the s pec ific mission, Minotaur vehicles can op erate from East and West Coast launch si tes
as shown in Figure 3.2-1. The corresponding range inclination capabilities are shown in Figure 3.2-2.
Specific Minotaur vehicle performance
parameters within those la u nc h inclination ranges
are presented in Section 3. 3. Per OSP-3 contract
requirements, baseline launch sites were
established and are shown in Table 3.2-1.

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Figure 3.1-1. Generic Minotaur IV Mission Profile

Minotaur IV Family of Launch Vehicles

Minotaur IV • V • VI User’s Guide Section 3.0 – General Performance

Figure 3.2-1. Flexible Processing and Portabl e GSE Allows Operations from Multiple Ranges or

Austere Site Options

Figure 3.2-2. Launch Site Inclinations

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3.2.1. Western Launch Sites

For missions requir ing h igh incl inati on orbi ts ( greater than 60°), launch es c an be conduc ted f rom fac ilities
at VAFB or Kodiak Island, AK, as shown in Figure 3.2-2. Inclinations below 72° from VAFB are possible,
but require an out-of-plane dogleg, thereby reducing payload capability. Minotaur IV is nominally
launched from the Californi a Space port f acilit y, Spac e Launc h Com plex 8 (SLC-8) operated by Spacepor t
Systems International ( SSI), on South V AFB. The laun ch fac ility at Kodiak Island, operated b y the Alask a
Aerospace Developm ent Corporation (AADC), can ac commodate the larger Minotaur V a nd VI vehicles
and has been used for both orbital and suborb ital lau nc hes including past launches of Minotaur IV.

3.2.2. Eastern Launch Sites

For easterly launch azimuths to achieve orbital inclinations between 28.5° and 55°, launches can be
conducted f rom fac ilities at Cape Cana veral Air Force Station , FL (CCAFS) or W allops I sland, VA (WFF).
Launches from Flor ida will nominally use the Space Florida launch f acilities at LC-46 on CCAFS which
can accommodate an y of the Minotaur vehicle co nfigurations. T ypical inclinations are from 28.5° to 50°;
however, higher inclination trajectories may require northerly ascent trajectories. These would need to
consider the potential of European overflight and require range safety assessment. The Mid-Atlantic
Regional Spaceport (MA RS) facilities at the W FF may be used for inclinations f rom 37.8° to 55°. Some
inclinations and/or altitudes may have reduced perfor mance due to range safety considerat ions and will
need to be evaluated on a case-by-case mission-specific basis.

3.2.3. Alternate Launch Sites

Other launch facilities can be readily used given the flexibility designed into the Minotaur IV vehicle,
ground support equipm ent, and the various interfaces. O rbital has experience launching ve hicles from a
variety of sites around the world. To meet the requirements of performing mission operations from
alternative, austere launch sites, Orbital can provide self contained, transportable shelters for launch
operations as an unpriced option. The m obile equivalent of the LCR is the Launch Support Van (LSV),
and the mobile LEV is the Launch Equipment Van.

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1

2

3.3. Performance Capability

Minotaur IV performanc e curves for circular orbits of various alt itudes and inclinations are shown on the
next several pages for launches from all four Spaceports in metric and English units. These perf ormance
curves provide the total m ass above the standard, no n-separating interface. The mass of the separation
system, and any Pa yload Adapter (PLA) that is attached to th e 38.81 in. interface, is to be accounted f or
in the payload mass allocat ion. T abl e 3.3-1 shows a number of c omm on optio ns and the mass associated
with each . Figures 3.3-1 and 3.3-2 show relative performance of the Minotaur IV family of launch vehicles
for representative launches from KLC and CCAFS.

Table 3.3-1. Common Mission Options and Associated Masses

(These Masses Must Be Subtracted from the LV Performance)

Option

Total Mass (kg)

(These Masses Must Be

Subtracted from the LV

Performance)

Portion of Total Mass

That Remains with SV

Post Separation (kg)

Enhanced Telemetry 9.85 0
TDRSS 8.54 0
62” Payload Adapter Cone
Two Piece Payload Adapter Cone (92” to 38”)
38” Orbital Separation Syst em

2

1

38” RUAG Low Shock Separation System (937S)

-10.32 0

9.07 0

12.24 4.0

19.89 6.16
38” RUAG Separation System (937B)2 18.25 5.18
38” Lightband
38” Softride and Ring

2

3

8.85 2.52

9 to 18 0

Notes: 1 For more information on these payload cone options, refer to Table 5.2.4.1-1.

2

For more information on these separation system options, refer to Table 5.2.5-1.

3

A range is provided for the softride option; actual mass is based on satellite requirements.

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Figure 3.3-1. Minotaur IV Fleet Comparis o n Performance Curves for SSO Out of KLC

Figure 3.3-2. Minotaur IV Fleet Comparis o n Performance Curves for

28.5° Inclination Orbits Out of CCAFS

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3.3.1. Minotaur IV LEO Orbits

Figure 3.3.1-1. Minotaur IV Performance Curves for VAFB Launches

Figure 3.3.1-2. Minotaur IV Performance Curves for KLC Launches

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Figure 3.3.1-3. Minotaur IV Performance Curves for CCAFS Launches

Figure 3.3.1-4. Minotaur IV Performance Curves for WFF Launches

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3.3.2. Minotaur IV+ LEO Orbits

Figure 3.3.2-1. Minotaur IV+ Performance Curves for VAFB Launches

Figure 3.3.2-2. Minotaur IV+ Performance Curves for KLC Launches

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Figure 3.3.2-3. Minotaur IV+ Performance Curves for CCAFS Launches

Figure 3.3.2-4. Minotaur IV+ Performance Curves for WFF Launches

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3.3.3. Minotaur VI LEO Orbits

Figure 3.3.3-1. Minotaur VI (92” Fairing) Performance Curves for CCAFS Launches

Figure 3.3.3-2. Minotaur VI (92” Fairing) Performance Curves for KLC Launches

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Figure 3.3.3-3. Minotaur VI (110” Fairing) Performance Curves for CCAFS Launches

Figure 3.3.3-4. Minotaur VI (110” Fairing) Performance Curves for KLC Launches

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3.3.4. Elliptical Orbits and High Energy Orbits

The Minotaur IV+, V, and VI+ are capable of supporting elliptical and high energy orbits, including
Geostationary Transf er Orbits (GT O), Medium Transfer Or bits (MTO), and Trans-Lunar Injec tion (T LI), as
shown in Figures 3. 3.4-1 through 3.3.4-5 and Tables 3.3.4-1 through 3.3.4-3. Orbital evaluates specific
high energy or elliptical missions on a case by case basis.

Figure 3.3.4-1. Minotaur IV+ Elliptical Orbits Performance Curve for KLC Launches

Figure 3.3.4-2. Minotaur V Elliptical Orbits Performance Curve for KLC Launches

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Figure 3.3.4-3. Minotaur VI+ Elliptical Orbits Performance Curves for KLC Launches

Figure 3.3.4-4. Minotaur V/VI+ High Energy Orbit Performance Curves for CCAFS Launches

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Figure 3.3.4-5. Minotaur V High Energy Orbit Performance Curve for WFF Launches

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GTO from CCAFS:

Payload Capability (Inclination 28.5°)

532 kg

866 kg

819 kg

MTO from CCAFS:

Payload Capability (Inclination 55°)
No Argument of Perigee Constraint

Payload Capability (Inclination 39°)

Argument of Perigee = 180°

Not Achievable

650 kg

Minotaur VI+

991 kg

1025 kg

Minotaur VI+

935 kg

988 kg

MTO from WFF:

Payload Capability (Inclination 55°)

Payload Capability (Inclination 39°)

603 kg

649 kg

Table 3.3.4-1. Geosynchronous Transfer Orbit (GTO) Performance For CCAFS

C3 = -16.3 km2/s2

Argument of Perigee = 180°

Inclination = 28.5°

Minotaur VI+ (110" Fairing)

Minotaur V

(92" Fairing)

(110" Fairing)

Vehicle

Minotaur V

Minotaur VI+ (92" Fairing)

Table 3.3.4-2. Medium Transfer Orbit (MTO) Performance For CCAFS

Vehicle

(Due to stage drops over land)

Table 3.3.4-3. Medium Transfer Orbit (MTO) Performance For WFF

C3 = -24.0 km2/s2

2185 lbm

2063 lbm

Argument of Perigee = 180

1173 lbm

1909 lbm

1806 lbm

1433 lbm

2261 lbm

2179 lbm

Vehicle

Minotaur V

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C3 = -24.0 km2/s2

No Argument of Perigee Constraint

1329 lbm

Argument of Perigee = 180°

1432 lbm

Minotaur IV • V • VI User’s Guide Section 3.0 – General Performance

Tolerance

Altitude

Stage 4 motor performance uncertainty and guidance

Altitude

Stage 4 motor performance and guidance algorithm

Altitude

Stage 4 motor performance and guidance algorithm

Guidance algorithm uncertainty and navigation

Table 3.5-1. Typical Pre-Separation Payload

Error Type

Angle

Rate

Yaw

±1.0°

≤1.0°/sec

Pitch

±1.0°

≤1.0°/sec

Roll

±1.0°

≤1.0°/sec

Spin Axis

±1.0°

≤10 rpm

Spin Rate

--

±3°/sec

3.4. Injection Accuracy

Minotaur IV injection accuracy limits are summarized in Table 3.4-1. Better accuracy can likely be
provided depending on spec ific mission charact eristics. For example, heavier pa yloads will t ypically have
better insertion accurac y, as will higher orbits. Furtherm ore, an enhanced option for increased ins ertion
accuracy is also available (Section 8.9) that utilizes the flight proven Hydrazine Auxiliary Propulsion
System (HAPS).

Table 3.4-1. Minotaur IV Injection Accuracy

Error Type

(Insertion Apse)

(Non-Insertion Apse)

(Mean)

Inclination ±0.2°

3.5. Payload Deployment

Following orbit insertion, the Minotaur IV
avionics subsystem can execute a series of
ACS maneuvers to provide the des ire d in itial
payload attitude prior to separation. This
capability may also be used to inc rem entally
reorient Stage 4 for the deployment of
multiple spacecraft with ind epen den t attitude
requirements. Either an inertially-fixed or
spin-stabilized attitude may be specified by
the customer. The maximum spin rate for a specific mission depends upon the spin axis moment of inertia
of the payload and the amount of ACS propellant needed for other attitude maneuvers. Table 3.5-1
provides the typical payload pointing and spin rate accuracies.

3.6. Payload Separation

Payload separation dynamics are highly dependent on the mass properties of the payload and the
particular separation s ystem utilized. The prim ary parameters to be considered are p ayload tip-off and
the overall separation velocity.

Payload tip-off refers to the angular velocity imparted to the payload upon separation due to payload
center-of-gravity (CG) offsets and an uneven distribution of torques and forces. Separation system
options are disc uss ed furth er in Section 5.2.4. O rbit al per form s a m iss ion-spe cif ic tip -off anal ysis for each
payload.

Separation velocities are d riven by the need to prevent recontac t b etween the pa yload and the Minotaur
final stage after separation. The value will typically be 0.6 to 0.9 m/sec (2 to 3 ft/sec).

(Worst Case)

±18.5 km (10 nmi)

±92.6 km (50 nmi)

±55.6 km (30 nmi)

Error Source

algorithm uncertainty

uncertainty and navigation (INS) error

uncertainty and navigation (INS) error

(INS) error

Pointing and Spin Rate Accuracies

3-Axis

Spinning

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3.7. Collision/Contamination Avoidance Maneuver

Following orbit insertion and payload separation, the Minotaur final stage will perform a
Collision/Contamination Avoidance Maneuver (C/CAM). The C/CAM minimizes both payload
contamination and the potential for recontact between Minotaur hardware and the separated payload.
Orbital will perform a recontact analysis for post-separation events.

A typical C/CAM begins shortly after payload separation. The launch vehicle performs a 90° yaw
maneuver designed to direct any remaining motor impulse in a direction which will increase the
separation distance bet ween the two bodies. After a delay to allow the distanc e between the spacecraft
and Stage 4 to incr ease to a safe level, the launc h vehicle begins a “crab-walk” maneuver to im part a
small amount of delta velocity, increasing the separation between the payload and the final stage.

Following the completion of the C/CAM maneuver as described above and any remaining maneuvers,
such as separating other small secondar y payloads or downlinking of delayed telemetry data, the ACS
valves are opened and the remaining nitrogen propell ant is expelled to meet Int ernational Space Debris
guidelines.

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4. PAYLOAD ENVIRONMENT CAUTION

The predicted environm ents provided in this user's guid e are for initial planning
purposes only.

Environments presented here bound typica l mission par ameters, but should not
be used in lieu of m ission-specific analyses. Mission-specific levels are pro vided
as a standard service and documented or referenced in the mission ICD.

This section provides det ails of the pr edicted enviro nmental c onditions the pa yload wil l experience dur ing
Minotaur ground operations, powered flight, and launch system on-orbit operations.

Minotaur ground operat ions include payload i ntegration and encaps ulation within the fair ing, subsequent
transportation to the la unch site and final vehicle inte gration activities. Powered f light begins at Stage 1
ignition and ends at final stage burnout. Minotaur on-orbit oper ations begin after final stage burnout and
end following payload separation. To more accurately define simultaneous loading and environmental
conditions, the powered flight portion of the mission is further subdivided into smaller time segments
bounded by critical, transient flight events such as motor ignition, stage separation, and transonic
crossover.

The environmental des ign and test criteria present ed have been derived using m easured data obtained
from many different sources, including Minotaur flights, Peacekeeper motor static fire tests, and other
Orbital system development tests and flights. The predicted levels presented are intended to be
representative of a standar d m iss ion and cont a in margins consistent with MIL-STD 1540B. Satel lit e mass,
geometry and structura l components var y greatly and will result i n significant differ ences from m ission to
mission.

Dynamic loading events that occur throughout various portions of the flight include steady-state
acceleration, transient low frequency acceleration, acoustic impingement, random vibration, and
pyrotechnic shock events.

4.1. Steady State and Transient Acceleration Loads

Design limit load factors due to the combined effects of steady state and low frequency transient
accelerations are largel y governe d by payloa d charact eristics . A mis sion-s pecif ic Coup led L oads An alysis
(CLA) will be perform ed, with c ustom er provided fin ite elem ent models of the pa yload, in or der to prov ide
precise load predictio ns. Results will be referenced in the m ission specific ICD. For preliminar y design
purposes, Orbital ca n provide initial Cen ter-of-Gravity (CG ) netloads given a pa yload’s mass properties,
CG location and bending frequencies.

4.1.1. Transient Loads

Transient events account f or approximatel y 90% of the total space vehicle loads with the remainder due
to steady state events . Transient loads are hig hly dependent on SV m ass, CG, natural frequ encies, and
moments of inertia as well as the chosen separation system and Payload Attach Fitting (PAF). All of
these were varied to devel op a range of transient lateral accelerati ons at the typical dominant transient
event and are shown as a function of payload mass in Figure 4.1.1-1 for Minotaur IV and Figure 4.1.1-2
for Minotaur IV+, V, VI, and VI+. These graphs cover a wide range of parameters whereas most
spacecraft/payloads will typic all y have later al acc elerations below 3.5 G’s.

Preliminary and final C LAs will be per formed f or each Minotaur m ission wher e the pa yload finite elem ent
model is coupled to the vehicle model. Forcing functions have been developed for all significant flight

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events and load cases. Results from the CLA are reported in the Acceleration Transformation Matrix
(ATM) and Load Transf ormation Matrix (LTM) as requested by the payload pro vider. A payload isolat ion
system is available as a non-standard option and is described in Section 8.10. This system has been
demonstrated to significantly reduce transient dynamic loads that occur during flight.

Figure 4.1.1-1. Payload CG Net Transient Lateral Acceleration (Minotaur IV)

Figure 4.1.1-2. Payload CG Net Transient Lateral Acceleration (Minotaur IV+, V, VI, and VI+)

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4.1.2. Steady-State Accelerat ion

Steady-state vehicle accelerations are determ ined from the vehicle rigid body analysis. Dra g, wind and
motor thrust are app lied to a vehicle m odel. A Monte-Carlo analysis is perform ed to determ ine variations
in vehicle accelerat ion due to changes in wi nds, motor perform ance and a erod ynamics . T he stead y-state
accelerations are added to transient accelerations from the CLA to determine the maximum expected
payload acceleration. Maximum steady state accelerations are dependent on the payload mass
enhancements chos en and vehicle config uration. Figure 4.1.2-1 depicts the maximum s teady state axial
acceleration as a function of payload mass. Lateral steady state accelerations are typically below 0.5 G’s.

Figure 4.1.2-1. Minotaur IV Family Maximum Axial Acceleration as a Function of Payload Mass

4.2. Payload Vibration Environment

The Minotaur payload v ibra tion en viro nments are low frequency random and sinusoida l vibr at io ns c r eated
by the solid rocket motor combustion processes and transmitted through the launch vehicle structure.
Additionally, higher f requency aeroacous tics energy is c reated by air flow over the surface of the vehicle.
Some of this aeroacoustic energy is transmitted via the launch vehicle structure to the payload. The
majority of the aeroaco ustic energy is transm itted to the payload d irectly as acoustic ener gy through the
fairing.

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4.2.1. Random Vibration

Payload random vibration is produced from two sources. The first is structural born from the launch
vehicle produced from motor burn and acoustics acting on the launch vehicle. This tends to be, low
frequency, less than 250 Hz, and c an be simulated using a base driven t est. The second source is fr om
acoustics acting o n the s pa c ecr af t. T his tends to be hi gh f r equency, greater than 250 Hz, and is not eas i l y
simulated using a base driven test. The response at the LV/S V interface is strongly dependent on the
unique spacecraft dynamics, including its response to the acoustic field. Therefore, structural born
random vibration environments are only defined up to 250 Hz and are shown in Figure 4.2.1-1.

Figure 4.2.1-1. Minotaur IV Family Payload Random Vibration Environment

Orbital recommends that the payload be subjec t to acoustic testing per Section 4.3, which will envelope
the high frequency (>25 0 H z) str uc tur al b or n random vibration, and that t h e payload be designed/qua lifie d
to meet the CLA results which envelope the low frequency (<250 Hz) structural born random vibration.

4.2.2. Sine Vibration

There are only two sourc es of s ine vibrat ion ex c it ati on on the Minotaur vehicle a n d they are defined at the
LV/SV interface as shown in Figures 4.2.2-1 and 4.2. 2 -2.

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Figure 4.2.2-1. Minotaur IV, IV+, and VI Payload Sine Vibration MPE Levels

Figure 4.2.2-2. Minotaur V and VI+ Payload Sine Vibration MPE Levels

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4.3. Payload Acoustic Environment

The acoustic environments to which the spacecraft will be exposed have been defined based on
measured acoustic d ata from previous flights which utilized the Peacek eeper Stage 1 motor and 92 in.
fairing. The data was adjusted to accou nt for differences in vehicle trajectories. The resulting acoustic
level, which also includes the damping of the acoustic blankets, is shown in Figure 4.3-1. Acoustic
environments for the optional 110” fairing are enveloped by these levels.

Figure 4.3-1. Minotaur IV Payload Acoustic Maximum Predicted Environment (MPE) with 1/3

Octave Breakpoints

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4.4. Payload Shock Environment

The maximum shock response spectrum at the base of the payload from the launch vehicle will not
exceed the flight limit levels (LV to Payload) in Figure 4.4-1 (Minotaur IV/IV+/VI) and Figure 4.4-2
(Minotaur V/VI+). For m issions that utili ze an Orb ital-supplied separation s ystem , the max imum expected
shock (LV to Pa yload) w ill be t he le ve l sho w n f or t h e c hos en s e par at io n s yst em. For missions t hat do not
utilize an Orbital-s upplied separation s ystem, the maximum expec ted shock (LV to Payloa d) is provided
and denoted as "Fairing Jettison Shock at Payload I/F".

For all missions, the shock response s pectrum at the base of the payload from payload even ts should not
exceed the levels in Figur e 4.4-3 (Pa yload to LV). Sh ock above this le vel could requ ire requalif ication of
launch vehicle components.

Figure 4.4-1. Minotaur IV Family Payload Shock Maximum Predicted Environment (MPE) –

Launch Vehicle to Payload

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Figure 4.4-2. Minotaur V/VI+ Payload Shock MPE – Launch Vehicle to Payload

Figure 4.4-3. Maximum Shock Environment - Payload to Launch Vehicle

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4.5. Payload Structural Integrity and Environments Verification

The spacecraft must possess sufficient structural characteristics to survive ground handling and flight
load conditions with margin in a manner that assures both safety and mission success.

Sufficient payload testing and/or analysis must be performed to show adequate margin to the
environments and loads s p ec if ied in Sec t ions 4.1 t hr o ugh 4. 4. The payload desig n s houl d comply with the
testing and design factors of saf ety as foun d in MI L-HNBK-340A (r ef. MIL-STD-1540B) and NASA GEVS
Rev. A June ’96. T he payload organizatio n must provide Orbital with verification via analyses and tests
that the payload can survive these environments prior to payload arrival at the integration facility.

4.6. Thermal and Humidity Environments

The thermal and hum idity environment to which the payload may be exposed during vehicle process ing
and pad operations are defined in the following sections.

4.6.1. Ground Operations

Upon encapsulation within the fairing and for the remainder of ground operations, the payload
environment will be maintained by a Heating, Ventilation and Air Conditioning (HVAC) Environmental
Control Unit (ECU). T he HVAC pro vides c ondition ed air to the pa yloa d in the PPF after fairing integrat ion.
HVAC is pro vided during transport, lifting operations, and at the launch pad. T he conditioned air enters
the fairing volum e at a location forward of the payload, exits aft of the pa yload and is pro vided up to t he
moment of launch. A diff user is designed into the air conditioning inlet to reduc e impingement velocit ies
on the payload. Class 10 K (ISO 7) can be pro vided in side a clean room and at the payload fair ing HVAC
inlet on a mission specific basis as an enhanced option.

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

a. Dry Bulb Temperature: 13 to 29 °C (55 to 85 °F) controllable to ±5 °C (±10 °F) of setpoint
b. Temperature environment lower limit is 55 °F (12.8 °C) due to the upper stage motor limits.
c. Standard Setpoint = 18.3 °C (65 °F)
d. Dew Point Temperature: 3 to 17 °C (38 to 62 °F)
e. Relative Humidity: determined by drybulb and dew point temperature selections and generally

controlled to within ±15%. Relative humidity is bound by the psychrometric chart and will be
controlled such that the dew point within the fairing is never reached.

f. Nominal Flow: 11.3 m

A diagram of the HVAC system is shown in Figure 4.6.1-1.

3

/min (400 cfm)

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Figure 4.6.1-1. Minotaur IV HVAC System Provides Conditioned Air to the Payload

4.6.2. Powered Flight

The maximum fairing inside wall temperature will be maintained at less than 93 °C (200 °F), with an
emissivity of 0.92 in the region of the payload. However, the payload will see significantly lower
temperatures and emissivity due to fairing
acoustic blankets. This temperature limit
envelopes the maximum temperature of any
component inside th e payload fairing with a vie w
factor to the payload.

The fairing peak vent rat e is typicall y less than 1.0
psi/sec, as shown in Figure 4.6.2-1. Fairing
deployment will be i nitiated at a time in flight t hat
the maximum dynamic pressure is less than 0.01
psf or the maximum free molecular heat ing rate is

2

less than 1136 W/m

(0.1 BTU/ft2/sec), as

required by the payload.

Figure 4.6.2-1. Typical Minotaur IV Fairing

Pressure Profile

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4.6.3. Nitrogen Purge (Non-Standard Service)

If required for spot cooling or purging of a payload com ponent, Orbital will provide GN2 f low to localized
regions in the fairing as a non-standard service. This option is discussed in more detail in Section 8.3.

4.7. Payload Contamination Control

All payload integrat ion procedures, and Orbital’s c ontamination control progr am have been designed to
minimize the payload’s exposure to contamination from the time the payload arrives at the payload
processing facility through orbit insertion and separation. The payload is fully encapsulated within the
fairing at the payload processing facility, assuring the payload environment stays clean in a Class
100,000 environment. La unch vehicle assemblies that aff ect cleanliness with in the encapsulat ed payload
volume include the f airin g a nd the payload cone assem bly. These assemblies are clean ed s uc h that t here
is no particulate or non-particulate matter visible to the norm al unaided eye when insp ected from 2 to 4
feet under 50 ft-candle inci dent light (Visibly Clean Level II). After encapsulation, the fairing envelope is
either sealed or maintained with a positive pressure, Class 100,000 (ISO 8) continuous purge of
conditioned air.

If required, the pa yload can be provided with enhanced contaminati on control as an option, providing a
Class 10,000 (ISO 7) env ironment, low outgassing, and Visibly Clea n Plus Ultraviolet cleanliness. With
the enhanced contam ination control option, th e Orbital-supplied elem ents will be cleaned and co ntrolled
to support a Class 10,000 clean room environment, as defined in ISO 14644-1 clean room standards
(ISO 7). This includes limiting volatile hydrocarbons to maintain hydrocarbon content at less than 15 ppm.

Also with the enhanced contamination control option, the ECU continuously purges the fairing volume
with clean filtered air and maintains humidity between 30 to 60 percent. Orbital’s ECU incorporates a
HEPA filter unit to provide ISO 7 (Class 10,000) air. Orbital m onitors the suppl y air for particu late matter
via a probe install ed upstream of the fairing inlet duct prior to connect ing the air source to the pa yload
fairing.

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(Optional)

Receive/
Transmit

L-Band
(L1/L2)

1575.42 /

1227.6

20.46 MHz

(P(Y) Code)

Spread

QPSK

Field Strength at

Cone

4.8. Payload Electromagnetic Environment

The payload Electromagnetic Environment (EME) results from two categories of emitters: Minotaur
onboard antennas and Range radar. All power, control and signal lines inside the payload fairing are
shielded and properly terminated to minimize the potential for Electromagnetic Interference (EMI). The
Minotaur payload fair ing is Radio Frequency (RF) opaque , which shields the payload from external RF
signals while the payload is encapsulated.

Table 4.8-1 lists the frequencies and maximum radiated signal levels from vehicle antennas that are
located near the payload d uring ground oper ations and po wered flight. T he specific EME exper ienced by
the payload during ground processing at the PPF and the launch site will depend somewhat on the
specific facilities that are utilized as well as operational details. However, typically the field strengths
experienced by the pa yload duri ng ground processin g with t he fairing in place ar e contro lled proce durall y
and will be less than 2 V/m from continuous sources and less than 10 V/m from pulse sources. The
highest EME during powered flight is created by the C-Band transpon der transmission, which results in
peak levels at the payload interface plane of 25.40 V/m at 5765 MHz. Range transm itters are typically
controlled to provi de a f ie ld s trength of 10 V/m or less i ns ide the fairing. This EM E s hou ld be c ompared to
the payload’s RF susceptibility levels (MIL-STD-461, RS03) to define margin.

Table 4.8-1. Minotaur IV Launch Vehicle RF Emitters and Receivers

SOURCE 1 2 3 4 5 6 7 8

Function

Band UHF C-Band C-Band S-Band S-Band S-Band S-Band
Frequency (MHz) 421 5765 5690 2240.5 2285.5 2260.5 2270.5
Bandwidth N/A 14 MHz 14 MHz 1.78 MHz 1.78 MHz 256 kHz 256 kHz

Power Output N/A 400 W (peak) N/A 10 W 10 W 5 W 5 W N/A
Sensitivity -107 dBm -70 dBm -70 dBm N/A N/A N/A N/A -123 dBm

Modulation Tone Pul se Code Pulse Code PCM/FM PCM/FM PCM/FM PCM/FM

Fwd Edge of the
Payload Adapter

Command

Destruct

Receive Transmit Receive Transmit Transmit Transmit Transmit Receive

N/A

Tracking

Transponder

1.99 V/m avg

(25.40 V/m per

0.5μs pulse)

Tracking

Transponder

N/A <11 V/m <11 V/m <6 V /m <6 V/m N/A

Launch
Vehicle

Enhanced

Instrumentation

Telemetry

GPB A GPB B GPB

Spectrum

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Minotaur IV • V • VI User’s Guide Section 5.0 – Payload Interfaces

Standard 92” Fairing with Standard 38” PAF

5. PAYLOAD INTERFACES

This section describes the available mechanical, electrical and Launch Support Equipment (LSE)
interfaces between the Minotaur launch vehicle and the payload.

5.1. Payload Fairing

5.1.1. 92” Standard Minotaur Fairing

Orbital’s flight proven 92-inch diameter payload fairing is used to encapsulate the payload, provide
protection and contam ination contr ol during gro und h andling, integr ation op erations and flight. T he fairing
is a bi-conic des ign made of graphi te/epox y face she ets with a luminum honeyco mb cor e. The two hal ves
of the fairing are structurally joined along their longitudinal interface using Orbital’s low contamination
frangible joint syst em. An additional circum ferential frangible joi nt at the base of the fair ing supports the
fairing loads. At separat ion, a gas press urization syste m is activated to press urize the fairin g deployment
thrusters. The fairing halves then rotate about external hinges that control the fairing deployment to
ensure that payload and launch vehicle clearances are maintained. All elements of the deployment
system have been demonstrated through numerous ground tests and flights.

5.1.1.1. 92” Fairing Payload Dynamic Design

Envelope

The fairing drawing in Fig ure 5.1.1.1-1 shows the
maximum dynamic envelope available in the
standard MIV configurat ion for the payload d uring
powered flight. The dynamic envelope shown
accounts for fairing and vehicle structural
deflections only. The payload contractor must
consider deflections du e to spacecr aft design and
manufacturing tolerance stack-up within the
dynamic envelope. Proposed payload dynamic
envelope violations must be approved by Orbital
via the ICD.

No part of the payload may extend aft of the
payload interface plane without specific Orbital
approval. Incursions below the payload interface
plane may be approved o n a case-by-case basis
after additional verification that the incursions do
not cause any detrimental effects. Vertices for
payload deflection must be given with the Finite
Element Model to evaluate payload dynamic
deflection with the C ouple d Loa ds Anal ysis (C LA).
The payload contractor should assume that the
interface plane is r igid; Orbital has accounted f or
deflections of the interface plane. The CLA will
provide final verific ation tha t the payload does not
violate the dynamic envelope.

Figure 5.1.1.1-1. Dynamic Envelope for

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110” Fairing with Standard 38” PAF

5.1.2. Optional 110” Fairing

A larger 110” diameter fairing design is available as an enhancem ent to accommodate payloads larg er
than those that can be fit in the stan dard 92” diamete r fairing. The larger fairing is prim arily intended for
use by Minotaur VI and VI+ pa yloads, with limited app lications ava ilable on other Minotaur configurations.
Flying the 110” fairing will result in approximately 200 kg performance impact and reduced launch
availability. The f airing, composite materials, structur al testing, separation and deplo yment systems are
similar to those of the heritage 92” f airing. The only appreciabl e change to the depl oyment system is the
use of a new thruster bracket that attaches to the boat-tail portion of the af t end of the f air ing. Deployment
margin is actuall y improved for the 110” fairi ng vs . the standard f air ing because the larg er diam eter of the
110” fairing draws the fairing mass radially outward and closer to the hinge pivot points.

Performance runs with the 110” fairing are included within Section 3.0.

5.1.2.1. 110” Fairing Payload Dynamic Design Envelope

Figure 5.1.2.1-1 shows the maximum dynamic envelope available in the larger 110” fairing for the payload
during powered flight. The dynamic envelope s hown a ccount s for fairing and vehicle s tructur al deflecti ons
only. The payload contractor must consider
deflections due to spacecraft design and
manufacturing tolerance stack-up within the
dynamic envelope. Proposed payload dynamic
envelope violations must be approved by Orbital
via the ICD.

No part of the payload may extend aft of the
payload interface plane without specific Orbital
approval. Incursions below the payload interface
plane may be approved on a case-by-case basis
after additional verification that the incursions do
not cause any detrimental effects. Vertices for
payload deflection must be given with the Finite
Element Model to evaluate payload dynamic
deflection with the C oupled Loads Anal ysis (CLA).
The payload contractor should assume that the
interface plane is rigid; Orbital has accounted for
deflections of the interface plane. The CLA will
provide final verificati on that the payload does not
violate the dynamic envelope.

Figure 5.1.2.1-1. Dynamic Envelope for Optional

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5.1.3. Payload Access Door

Orbital provides one 457 mm by 610 mm (18 in. b y 24 in.) payload fairin g access door to provide acces s
to the payload after f airing mate. The door can be pos itioned according to payload re quirements within
the cylindrical sectio n of the f airing, pr oviding acc ess to the p aylo ad witho ut hav ing to rem ove an y porti on
of the fairing or break electrical connections. If necessary, the fairing acc ess door may be place d within
the lower conic sec tion of the f airing, ho wever th e sta ndard s ize is reduc ed t o 356 mm by 559 mm (14 in.
by 22 in.). The specific location is defined and controlled in the payload ICD. See Figure 5.1.3-1 for
available Access D oor loca tions. Addi tional acc ess do ors can r eadily b e provide d as a n enhanc ed optio n
(see Section 8.4).

Figure 5.1.3-1. Available Fairing Access Door Locations

5.2. Payload Mechanical Interface and Separation System

Minotaur provides for a s tandard non-separat ing payload interf ace. Orbital will provi de all flight hardware
and integration services necessary to attach non-separating and separating payloads to the Minotaur
launch vehicle. Payload ground handling equipment is typically the responsibility of the payload
contractor. All attachm ent hardware, whet her Orbital or c ustomer provided, m ust contain lock ing features
consisting of locking nuts, inserts or fasteners. Additional mechanical interface diameters and
configurations can readily be provided as an enhanced option.

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Minotaur IV • V • VI User’s Guide Section 5.0 – Payload Interfaces

Figure 5.2.1-1. Minotaur IV Coordinate System

5.2.1. Minotaur Coordinate System

The Minotaur IV Launch Vehicle coordinate s ystem is defined in F igure 5.2.1-1. F or clocking refer ences,

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degree marks are counterclockwise when forward looking aft. The positive X-axis is forward along the
vehicle longitudina l centerl ine, the positive Z axis is along the 180 deg angu lar, a nd the pos itive Y axis is
along the 90 deg angular s tation, and completes the orthogonal system . The origin of the LV coordin ate
system is centered at the Stage 1 nozzle exit plane of the LV and the vehicle center line (X = 0.0 in., Y =

0.0 in., Z = 0.0 in.).

5.2.2. Orbital Supplied Mechanical Interface Control Drawing

Orbital will provide a toleranced Mechanical Interface Control Drawing (MICD) to the payload contractor to
allow accurate machining of the fastener holes. The Orbital provided MICD is the only approved
documentation for drilling the payload interface.

5.2.3. Standard Non-Separating Mechanical Interface

Orbital’s payload interface design provides a
standard interface that will accommodate m ultiple
payload configurations. The Minotaur IV baseline is
for payloads to provide their own separation
system or for payloads that will no t separate. The
standard interface is a 986 mm (38.81 in.) diameter
bolted interface. A butt joint with 60 holes
(0.281 in. diameter) designed for ¼ in. fasteners is
the payload mounting sur face as shown in Figure

5.2.3-1.

5.2.4. Optional Mechanical Interfaces

Alternate or multiple payloa d configura tions can be
accommodated with th e use of a variety of payload
adapter fittings as listed in Table 5.2.4-1. The
Minotaur IV family of Launch Vehicles allows
flexibility in mounting patterns and configurations
(Figure 5.2.4-1).

Figure 5.2.3-1. Standard, Non-separating 38.81”

Diameter Payload Mechanical Interface

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Table 5.2.4-1. Minotaur IV Payload Adapter Fitting Options

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Figure 5.2.4-1. Optional, Non-Separating 62.01” Diameter Payload Mechanical Interface

5.2.4.1. Payload Cone Interfaces

Several different payload cones can be provided to meet mission unique interface requirements. The
baseline Minotaur IV 38 .81 in. payload interf ace is desc ribed in Section 5.2.3. Ho wever, Orbita l has other
flight proven payload opti ons. One option maintains the 986 mm (38.81 in.) inte rface, but increases the
amount of fairing volum e by using a two piece paylo ad cone that moves the interface approx imately 203
mm (8 in.) aft. T his option adds 9 kg to the L V. Orbital c an also provide other opt ions, such as 1194 mm
(47 in.) or 1575 mm ( 62 in.) interf aces required b y some separ atio n system s . These options are sho wn in
Table 5.2.4-1 with corresponding fairing envelopes shown in Figures 5.2.4.1-1 through 5.2.4.1-3.

5.2.4.2. Dual and Multi Payload Adapter Fittings

The Minotaur launch vehicle design f lex ibilit y and p erf orm ance r eadily ac comm odates m ultiple spacec raf t
that are independently deployed when required.

5.2.4.2.1. Dual-Payload Adapter Fitting

Provisions for larger multiple payloads exist for the Minotaur IV launch vehicle. A flight proven Dual
Payload Attach Fitting (DPAF) supports delivery of two primary spacecraft to orbit. The structure that
supports the dual pa yload configuration includes a 1600 mm (63 in.) diam eter cylindrical section that is
configurable in height depending on payload unique requirements. In the DPAF configuration, the aft
positioned spacecr aft m ounts to a co ne inside the c ylinder which is in turn m ounted to t he forward flange
of the 62 in. payload adapter cone. The for ward positioned spacec raft is then mounted t o a cone on the
forward end of the DPAF cylinder using a 986 mm (38.81 in.) separation system. After the forward

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Figure 5.2.4.1-1. Dynamic Envelope for

Standard 92” Fairing with

Optional 38” 2-Piece Payload Cone

positioned spacecraf t deploys, its res pective payload c one is separated fr om the launch veh icle followed
by deployment of the aft spacecraft from inside the DPAF cylinder, also using a 986 mm (38.81 in.)
separation system . The separation system s are addressed in Section 5.2.5. The DPAF is qua lified to a
maximum height of 2.26 m (89 in.). Both payloads would interface to the stand ard, non-separating 986
mm (38.81 in.) diameter mechanical interface shown in 5.2.3-1. The fairing envelope with the DPAF
option is shown in Figure 5.2.4.2.1-1.

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Figure 5.2.4.1-2. Dynamic Envelope for

Standard 92” Fairing with

Optional 62” Payload Cone

Minotaur IV • V • VI User’s Guide Section 5.0 – Payload Interfaces

Figure 5.2.4.1-3. Dynamic Envelope for

Optional 110” Fairing with

Optional 62” Payload Cone

5.2.4.2.2. Multi-Payload Adapter Fitting (MPAF)

The Multi-Payload Adapter Fitting (MPAF) utilizes a flight proven multi-payload design. The MPAF
supports up to eight indi vidual payloads includ ing four ESPA class (610 b y 711 by 965 mm (24 by 28 x
38 in.) envelope), 181 kg (400 lbm) payloads on the Multiple Payload Adapter Plate (MPAP) and four
secondary 29.5 kg ( 65 lbm) payloads o n the adapter c ylinder with an allowable size envelope of 483 by
495 by 1219 mm (19 by 19.5 by 48 in.) each. The ada p ter cylinder can also accommodate two 59 kg (130
lbm) payloads in place of four 29.5 kg (65 lbm) payloads. The upper MPAP plate can also be
implemented independ ent of the ad apt er cylinder described a bo ve t hat al lo ws a s ingle M ino taur I V lau nch
vehicle to support f our Evolved Expendable Launch Vehic le (EELV) Secondar y Payload Adapter ( ESPA)
class payloads. The fairing envelope with the MPAF option is shown in Figure 5.2.4.2.2-1. The
mechanical interface to the MPAP is shown in Figure 5.2.4.2.2-2.

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Figure 5.2.4.2.1-1. Dynamic Envelope for

Standard 92” Fairing with

Optional DPAF

Minotaur IV • V • VI User’s Guide Section 5.0 – Payload Interfaces

Figure

Payload Adapter

Figure 5.2.4.2.2-1. Dynamic Envelope for
Standard 92” Fairing with Optional MPAF

5.2.4.2.3. Minotaur V and VI+ Payload Adapter Fitting

The Minotaur V and VI+ baseline i nterfac e is the stan dard 38. 81 in. non-separat ing inter face as s hown in
Figure 5.2.3-1. The Minot aur V and VI+ fair ing envelope with this interf ace is show in Figure 5.2.4.2.3-1.
In addition to the b as eline i nter f ac e, there is an optional Min ota ur V Payload Attach Fitting ( PAF) bet w een
the LV uppermos t stage and payload. It is an anisogrid structure constr ucted of a graphite epox y lattice
winding that attaches t o th e forward f lange of th e upp ermos t stage forward c ylinder . The fair ing envelo pe
with this PAF installed is sho wn in Figure 5.2.4.2.3-2. The PAF adapts to a 803 mm (31.6 in.) diameter
spacecraft interface ring, as shown in Figure 5.2.4.2.3-3.

5.2.5. Optional Separation Systems

Three separation system options are offered as flight proven enhancements for Minotaur IV family of
launch vehicles. All systems are configurable to various interface diameters and have extensive flight
history. These separa tion systems include the Or bital Pegasus-developed marmon clamp band system ,
Planetary Systems Corp. Motori zed Li ghtban d (MLB) System , and RU AG lo w-sh ock marm on c lamp band
system. Through this enhanc ement, Orbital proc ures the qualifie d separation syst em hardware, conducts

5.2.4.2.2-2. Optional Multi-

Plate (MPAP) Non-Separating Mechanical

Interface – Accommodates 2 to 4 ESPA-Class

Payloads

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Figure 5.2.4.2.3-1. Dynamic Envelope for

Standard 92” Fairing and Minotaur V / VI+

Enhanced Performance Option

separation testing and analyses, and integrates the system onto the launch vehicle. The separation
system options are summarized in Table 5.2.5-1.

The primary separatio n parameters associated with a separation system are payload tip-of f and overall
separation velocit y. Pa yload tip-of f ref ers to the ang ular veloc ity im parted to the pa yload upon s eparat ion
due to payload CG offs ets and an un even d istribu tion of torques and for ces. Pa yload t ip-of f r ates induced
by the separation system s presented are generally under 1 deg/ sec per axis. Entering into the pa yload
separation phase, the laun ch vehicle reduces vehicle rates . The combined tip-off rate of the separ ation
system and launch ve hicle is generally less than 2 deg/sec about each axis when spacecraf t mass CG
offsets are within specified limits presented in Sect ion 5.4.1. Separation velocities are usua lly optimized to
provide the spacecraft with the lowest separation velocity while ensuring recontact does not occur
between the payload and the Minotaur upper stage after separation. The spacecraft is ejected by
matched push-off springs with sufficient energy to produce the required relative separation velocity to
prevent re-contact with the spacecraft. If non-standard separation velocities are needed, alternative
springs may be substituted on a mission-specific basis as a non-standard service. Payload separation

Figure 5.2.4.2.3-2. Dynamic Envelope for

Standard 92” Fairing and Minotaur V / VI+

Enhanced Performance Option

with Optional PAF

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dynamics are highly dependent on the mass
properties of the payload and the particular
separation system utilized. Typical separation
velocity is 0.6 to 0.9 m/sec (2 to 3 ft/sec). As a
standard service, Orbital performs a mission­specific tip-off and separation analyses for each
spacecraft.

5.2.5.1. Orbital 38” Separatio n S yst em

The flight proven Orbital 38” separation system,
Figure 5.2.5.1-1, is more suitable for lighter weig ht
payloads and is com pos ed of two ri ngs co nnect ed
by a marmon clamp band which is separated b y
redundant bolt cutters. This system has flown
successfully on over forty Orbital launch vehicle
missions to date. The weight of hardware
separated with the payload is approximately 8.7
lbm (4.0 kg). Orbital-provided attachment bolts to this interface can be inserted from either the launch
vehicle or the payload side of the interface via the through-holes in the separation system flange.

Table 5.2.5-1. Minotaur IV Separation System Options

Figure 5.2.4.2.3-3. Minotaur V PAF Non-

Separating Mechanical Interface

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Figure 5.2.5.1-1. Orbital 38” Separation System

5.2.5.2. Planetary Systems Motorized Lightband (MLB)

The Planetary System s MLB, Figure 5.2.5.2-1, prov ides a fully qualified and flig ht proven low shock and
lightweight option f or us e on Min ota ur missions. Multiple sizes of MLBs have previously flown on Minotaur
vehicles. The MLB uses a s ystem of mechanical ly-act uated h inged leaves, s prin gs, and a dual redunda nt
release motor to sep arate the upper ring (mounted t o the spacecraft) from the lower ring. The MLB is
flexible and configurable to support various separation force requirements and number of required
separation connector s. The MLB upper ring inter faces to the spacecraft thr ough holes in the upper rin g
and remains attach ed after separat ion adding appr oximately 2.04 k g (4.5 lb) of m ass. Due to the uniq ue
design of the system and space constraints for toolin g, Orbital provided soc ket head cap screw mating
hardware must be inserted from the launch vehicle side. The MLB offers the unique abilit y to perform
separation verification tests both at a component and system level.

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5.2.5.3. RUAG 937 Separation Systems

There are two available RUAG separation
systems. The traditional RUAG 937B is a lower
cost, flight proven design composed of two rings
and a clamp band separ ated by bolt cutters. The
RUAG 937S separatio n system, Figure 5.2.5.3-1,
is flight proven, low-shock separation s ystem that
offers outstanding load capability. This s ystem is
composed of two rings and a clamp band
separated by a Clamp Band Opening Device
(CBOD) rather than traditional bolt cutters. The
CBOD uses a redundant, ordnance initiated pin
puller device to con vert strain energy, created by
the clamp band tension, into kinetic energy
through a controlled event that greatly reduces
separation shock. Hardware separated with the
payload is approximately 5.18 kg (11.40 lbm) for
the 937B and 6.2 kg (13.6 lbm) for the 937S.
Orbital-provided attac hment bolts to this interface
can be inserted from either the launch vehicle or
the payload side of the interface.

5.3. Payload Electrical Interfaces

The payload electrical interface supports battery
charging, external power, discrete commands,
discrete telemetry, analog telemetry, serial
communication, payload separation indications,
and up to 16 separate ordnanc e discretes. If an optiona l Orbital-provided separation s ystem is utilized,
Orbital will provide all the wiring t hrough the separa ble interface plane. If the option is not exercised the
customer will be responsible to provide the wiring from the spacecraft to the separation plane.

5.3.1. Payload Umbilical Interfaces

Two dedicated payload umbilicals are provided with 60 circ uits each from the ground to the spacecraft.
These umbilicals are dedicated pass through harnesses f or ground processing support. They allow the
payload command, control, monitor, and power to be easily configured per each individual user’s
requirements. The umbilical wiring is configured as a one-to-one from the Payload/Minotaur interface
through to the payload EGSE i nterface in the Launch Equipment Vault, th e closest location f or operating
customer supplied payload EGSE equipment. The length of the internal umbilicals is approximately

7.62 m (25 ft). T he length of the external um bilicals from the LEV/SEB t o the launch vehicle ranges from

approximately 38.1 m (125 ft) to 99.1 m (325 ft) depending on the launch site chosen for the mission.

Figure 5.3.1-1 details the pin outs for the standard interface umbilical. All wires are twisted, shielded
pairs, and pass throug h the entire umbilical system , both vehicle and ground , as one-to-one to sim plify
and standardize the payload umbilical configuratio n requirements while providin g maximum operational
flexibility to the payload provider.

Figure 5.2.5.2-1. 38” Planetary Sciences

Motorized Lightband

Figure 5.2.5.3-1. RUAG 937S 38”

Separation System

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Figure 5.3.1-1. Payload 1:1 Umbilical Pin Outs

5.3.2. Payload Interface Circuitry

Standard interface circ uitr y passing thr ough the pa yload-to-launch v ehicle e lectric al conn ectio ns is shown
in Figure 5.3.2-1. This figure details the interface cha racteristics for launch vehic le commands, discrete
and analog telem etry, sepa ration lo opbac k s, pyro initiation, and s erial com m unications interf aces with the
launch vehicle avionics systems.

5.3.3. Payload Battery Charging

Orbital provides the capability for remote controlled charging of payload batteries, using a customer
provided batter y charger. This power is routed through the payload umbilica l cable. Up to 5.0 amper es
per wire pair can be acco mmodated. The payload batter y charger should be sized to withstand the line
loss from the LEV to the spacecraft.

Figure 5.3.2-1. Payload Electrical Interface Block Diagram

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5.3.4. Payload Command and Control

The Minotaur standard interface provides discrete sequencing commands generated by the launch
vehicle’s Ordnance Driver Module ( ODM) that are available to the payload as closed circu it opto-isolator
command pulses of 5 A in lengths of 35 ms minim um. The total n umber of ODM disc retes is s ixteen ( 16)
and can be used for any combination of (redundant) ordnance events and/or discrete commands
depending on the payload requirements.

5.3.5. Pyrotechnic Initiation Signals

Orbital provides the capability to directly initiate 16 separate pyrotechnic conductors through two
dedicated MACH Or dna nc e Dr i ver M od ules ( O D M) . Ea ch O DM pr ov id es for up to eight driv ers c apa bl e of
a 5 A, 100 ms, current limited pulse into a 1.5 ohm resistive load. All eight channels can be fired
simultaneously with an accuracy of 1 ms between channels. In addition, the ODM channels can be
utilized to trigger high im pedance discrete events if requir ed. Safing for all pa yload ordnance events will
be accomplished either through an Arm/Disarm (A/D) Switch or Safe Plugs.

5.3.6. Payload Telemetry

The baseline telemetr y subsystem capability provides a num ber of dedicated payload discrete (bi-level)
and analog telemetr y monitors through ded icated chan nels in the ve hicle enco der. Up to 24 c hannels will
be provided with t ype and data rate being def ined in the mission requir ements document. The pa yload
serial and analog data will be embedded in the baseline vehic le telemetry format. For disc rete monitors,
the payload custom er must provide the 5 Vdc source and the return path. T he current at the payload
interface must be less than 10 mA. Separation breakwire monitors can be specified if required. The
number of analog channel s available for payload telem etry monitoring is dependent on the frequency of
the data. Payload telem etry requirements and signal characteristics will be spec ified in the Payload ICD
and should not change once the final telemetr y format is released at approx imately L-6 months. Orbital
will record, archive, and reduce the data into a digital format for delivery to the payloaders for review.

Due to the use of s tr ateg ic as s ets , Mi not aur IV telemetr y is subj ec t t o the provisions of t he Str ateg ic Arm s
Reduction Treaty (START). ST ART treaty provisions require that c ertain Minotaur vehicle telemetry be
unencrypted and provide d to the START treaty office for dissemination to the s ignatories of the treaty.
The extent to which START applies to the payload telemetry will be determined by SDL. Encrypted
payload telemetry can be added as a non-standard service pending approval by SDL and the START
treaty office.

5.3.7. Payload Separation Monitor Loopbacks

Separation breakwire monitors are required on both sides of the payload separation plane. With the
Orbital provided separation systems, Orbital provides three (3) separation loopbacks on the launch
vehicle side of the separation plane for positive payload separation indication.

The payload will pro vide t wo (2) sep aration loop back circuits on t he pa yload s ide of the s eparat ion p lane.
These are typicall y wired int o diff erent s eparat ion conn ector s for redund ancy. T hese br eak wires are used
for positive separation indication telemetry and initiation of the C/CAM maneuver.

5.3.8. Telemetry Interfaces

The standard Minotaur payload i nterface pro vides a 1 6 Kbps RS-422/RS-485 serial i nterface for payload
use with the flexibility to support a variety of channel/bit rate requirements, and provide signal

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Table 5.4.2-1. Payload Mass Properties

Component

Accuracy

Mass

±0.5%

Principal Moments of Inertia

±5%

Cross Products of Inertia

±2.7 kg – m2

(±2.0 slug – ft2 )

Center of Gravity X, Y,

and Z Axes

±0.64 cm

(±0.25 in.)

conditioning, PCM f ormatting (programm able) and data trans mission bit rates. T he number of channels ,
sample rates, etc. will be defined in the Payload ICD.

5.3.9. Non-Standard Electrical Interfaces

Non-standard services such as serial command and telemetry interfaces can be negotiated between
Orbital and t he pa yload on a mission-by-mission basis. The s election of the separ atio n s ystem could a lso
impact the payload interface design and will be defined in the Payload ICD.

5.3.10. Electrical Launch Support Equipment

Orbital will pr ovide space for a rack of custom er supplied EGSE in the LCR, and at the on-pad LEV or
SEB. The equipment will interface with the launch vehicle/spacecraft through either the dedicated payload
umbilical interface or direc tly through the payload acc ess door. The payload cus tomer is responsible for
providing cabling f or their EGSE with in the LCR, LEV, and SE B to the appr opr iate um bilical inter f ace.

Separate payload ground processing harnesses that mate directly with the payload can be
accommodated through the payload access door(s) as defined in the Payload ICD. The payload will
provide all cabling for this operation.

5.4. Payload Design Constraints

The following sections provide design constraints to ensure payload compatibility with the Minotaur
launch vehicle.

5.4.1. Payload Center of Mass Constraints

Along the Y and Z-axes, the payload CG must be within 1.0 inch (2.54 cm) of the vehicle centerline.
Payloads whose CG exten d be yond the 1.0 inc h lat eral off set lim it will req uire Orbita l to verif y the spec ific
offsets that can be accommodated.

5.4.2. Final Mass Properties Accuracy

In general, the final mass properties statement
must specify payload weight to an accuracy of
±0.5% of the payload m ass, the center of gravity
to an accuracy of at least 0.64 cm (0.25 in.) in
each axis, moment of inertia to ±5%, and the
products of inertia to an accuracy of less than

2.7 kg-m

1. However these accuracies may vary on a

mission specific bas is. In additi on, if the pa yload
uses liquid propellant, the slosh frequency must be provided to an accuracy of 0.2 Hz, along with a
summary of the method used to determine slosh frequency.

5.4.3. Pre-Launch Electrical Constraints

Prior to launch, all payloa d electrical interface c ircuits are constrained to ensur e there is no current flow
greater than 10 mA across the payload electrical interface plane. The primary support structure of the
spacecraft shall be electrically conductive to establish a single point elec tr ic al gro und.

5.4.4. Payload EMI/EMC Constraints

2

(2.0 slug-ft2), as shown in Table 5.4.2-

Measurement Tolerance

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The Minotaur avion ics ar e in close prox im ity to the pa yload inside the f airi ng such that r adiate d em issions
compatibility is paramount. Orbital places no firm radiated emissions lim its on the payload oth er than the
prohibition against RF transm issions within t he payloa d fairing. Pr ior to launch , Orbital re quires rev iew of
the payload radiated emission levels (MIL-STD-461, RE02) to verify overall launch vehicle EMI safety
margin (emission) in accordance with MIL-E-6051. Payload RF transmissions are not permitted after
fairing mate and prior to an ICD specified time after separation of the payload. An EMI/EMC analysis may
be required to ensure RF compatibility.

Payload RF transm ission freque ncies m ust be coord inated wit h Orbital and range off icials to ensur e non­interference with Minot a ur and r an ge tr a nsmissions. Additionally, the customer must sc hedule al l RF tes ts
at the integration site with Orbital in order to obtain proper range clearances and protection.

5.4.5. Payload Dynamic Frequencies

To avoid dynamic c oupling of the payload m odes with the natural frequency of the launch vehicle, the
spacecraft should be d esigned with a struc tural stif fness to ens ure that t he later al fundam ental freque ncy
of the spacecraft, f ixed at the spacecraft interface is typically greater than 15 H z lateral. However, this
value is significantly affected by other factors such as incorporation of a spacecraft isolation system
and/or separation s ystem . Theref ore, the fin al determ ination of com patibility m ust be m ade on a miss ion­specific basis.

5.4.6. Payload Propellant Slosh

A slosh model should be provi ded to Orbital in either the pendu lum or spring-mass form at. Data on first
sloshing mode are required and data on higher order modes are desirable. Additional critical model
parameters will be established during the mission development process. The slosh model should be
provided with the payload finite element model submittals.

5.4.7. Payload-Supplied Separation Systems

If the payload employs a no n-Orbital separation s ystem, t hen th e shoc k delivered to the LV interface must
not exceed the limit level char acterized in Secti on 4.3 (Figure 4.4-2). Shock above the stated level could
require a requalification of LV components.

5.4.8. System Safety Constraints

OSP considers the saf ety of personnel an d equipment to be of paramount importance. AFS PCM 91-710
outlines the safet y design criteria for Minot aur payloads. These ar e compliance docum ents and must be
strictly followed. It is the re sponsibilit y of the custom er to ensure that the payloa d meets all OSP, O rbital,
and range imposed safety standards.

Customers designing pa yloads that employ hazardo us subsystems are advised to contact OSP early in
the design process to verify compliance with system safety standards.

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6. MISSION INTEGRATION

6.1. Mission Management Approach

OSP-3 is m anaged t hrough U.S. Air F orce, Space an d Missi le S ystems Center , Spac e Developm ent an d
Test Directorate (SD) Launch Systems Division ( SDL). SD/SDL ser ves as the prim ary point of contac t for
the payload customers for the Minotaur launch service. The organizations involved in the Mission
Integration Team are shown in Figure 6.1-1. Open communication between SD/SDL, Orbital, and the
customer, with an emphasis on timely data transfer and prudent decision-making, ensures efficient launch
vehicle/payload integration operations.

Figure 6.1-1. Mission Integration Team

6.1.1. SD/SDL Mission Responsibilities

SD/SDL is the pr imary focal point for all c ontractual and tec hnical coordination. SD/SDL c ontracts with
Orbital to provide the L aunc h Vehicle, launch integrat ion, and comm ercial f acilities (i.e. spacepor ts, clean
rooms, etc.). Separately, they contract with Government Launch Ranges for launch site facilities and
services. Once a m ission is identifie d, SD/SDL wil l assign a government Mission Manag er to coordinate
all mission planning and contracting activities. SD/SDL is supported by associate contractors for both
technical and logist ical support, capitalizing on their extensive expertise and background knowledge of
the Peacekeeper booster and subsystems.

6.1.2. Orbital Mission Responsibilities

As the launch vehicle provider, Orbital’s responsibilities fall into four primary areas:

a. Launch Vehicle Program Management
b. Mission Management
c. Engineering
d. Launch Site Operations

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The Minotaur organization uses highly skilled personnel with extensive Minotaur experience. The
Minotaur program is led by a Progr am Director who reports directly to Orbital ’s Launch Systems Group
General Manager and has full responsibility for mission success. This direct line to executive
management provides high visibility, ensuri ng access to critical org anizational res ources. Supporting the
Program Director is the Minotaur Chief Engineer, who provides technical direction and oversight to
maintain standard practices across Orbital’s family of Minotaur launch vehicles.

For new missions, a Pr ogr a m Management team is as s igne d. L ead ing th is team is the Program Manag er,
whose primar y responsibilit ies include developing staff requirements, inter preting contract requirements
as well as managin g schedules and budg ets for the mission. A Program Engineering Mana ger (PEM) is
assigned to provide management and technical direction to all engineering department personnel
assigned to the mission. The PEM is th e singl e foc al point f or all engineer ing activ ity an d func tions as the
chief technical lead f or the mission and technical advis or to the Program Manag er. In addition, the PEM
serves as the single point of contact for the OSP-3 Government COR.

Orbital also assigns a Mission Manager that serves as the primary interface to the SD/SDL Mission
Manager and payload provider. This person has overall mission responsibility to ensure that payload
requirements are met and that the appropriate launch vehicle services are provided. They do so via
detailed mission pla nning, payload integr ation scheduling, s ystems engineering, m ission-peculiar design
and analyses coordination, payload interface definition, and launch range coordination. The Orbital
Mission Manager will jointly chair Working Group meetings with the SD/SDL Mission Manager.

Engineering Leads and their supporting engineers conduct detailed mission design and analyses, perform
integration and test ac tivities, and f ollow the har dware to the f ield site to ens ure continu ity and m aximum
experience with that mission’s hardware.

Launch Site Operations are carried out by the collective Minotaur team as detailed in Section 7.0. A
Launch Site Integrati on and Operations lead are t ypically assigned an d on-site full-tim e to manage day­to-day launch site activities.

6.2. Mission Planning and Development

Orbital will assist th e customer with miss ion planning and development associated with Minotaur launch
vehicle systems. These services include interface design and configuration control, development of
integration processes, launch vehicle analyses and facilities planning. In addition, launch campaign
planning that includes range services, integrated schedules and special operations.

The procurement, anal ysis, inte gration an d test ac tiv ities requir ed to p lace a cus tom er’s payload into or bit
are typically conducted over a 26 month standard sequence of events called the Mission Cycle. This
cycle normally begins 24 months before launch, and extends to 8 weeks after launch.

The Mission Cycle is initiated upon receipt of the contract authority to proceed. The contract option
designates the pa yload, launch date, and bas ic mission param eters. In response, th e Minotaur Program
Manager designates an Orbital Mission Manager who ensures that the launch service is supplied
efficiently, reliably, and on-schedule.

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The typical Mission Cycle interweaves the following activities:

a. Mission management, docum ent exc hanges , m eetings, and formal reviews requir e d to c oordi nate

and manage the launch service.
b. Mission analyses and payload integration, document exchanges, and meetings.
c. Design, review, procurement, testing and integration of all mission-peculiar hardware and

software.
d. Range interface, safety, and flight operations activities, document exchanges, meetings and

reviews.

Figure 6.2-1 details the typ ical Mission Cycle and how this cycle folds into the Orbital vehic le production
schedule with typical pa yload activities and m ilestones. A typical Mission C ycle is based on a 24 month
interval between m ission authorization and launc h. This interval ref lects the OSP-3 contrac tual schedule
and has been shown to be an eff icient schedule based on Orbital’s past program execution experience.
OSP-3 does allow flexibility to negotiate either accelerated or extended mission cycles that may be
required by unique payload requirem ents. Payload sce narios that m ight drive a change in th e duration of
the mission cycle incl ude those that have funding lim itations, rapid response dem onstrations, extensive
analysis needs or contain highly complex payload-to-launch vehic le int egrat ed designs or tests.

Figure 6.2-1. Typical Mission Integration Schedule

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A typical miss ion field integration schedule is provided in Figure 6.2-2. The f ield integration schedule is
adjusted as required based on the mission requirements, launch vehicle configuration and launch site
selection.

Figure 6.2-2. Typical Mission Field Integration Schedule

6.2.1. Mission Assurance

The OSP-3 contract has three tailored levels of Mission Assurance (MA); Category 1, Category 2 and
Category 3. These categori es pr ovide pr ogr ess i ve ly in c reas ing le vels of governm ent ov er sigh t, above and
beyond Orbital rigorous internal MA standards.

Category 1 MA is the simplest, relying on O r b ita l's rob us t int er na l M A sta ndar ds a nd proc es s es, a nd does
not required SMC flight worthiness certification or Government IV&V oversight . Category 1 missions wil l
be licensed under Federal Aviation Administration (FAA) licensing guidelines.

Category 2 MA builds upon Category 2 and dictates that Orbital provide additional information and
support for the government's MA efforts and the government's Independent Readiness Review Team
(IRRT). Orbital will provide support for SMC's Spaceflight Worthiness Certification, independent IV&V,
requirements decom position and verificatio n, testing (planning, qual ification, design verific ation), as well
as additional reviews and activities both pre and post launch. Category 2 MA represents what has
traditionally been the standard level of MA on past Minotaur missions.

Category 3 MA builds upon the requirements of Category 2 and are subject to increased breadth and
depth of governm ent IV&V and insight. Up t o ten dedi cated IRRT reviews m ay be require d, with m onthly
1-day Program Management Reviews throughout the period of performance, as well as weekly 2-hour
telecons to comm unicate current status of concer ns and action items. C ategory 3 is intended mainly for
high value DoD missions similar to Acquisition Category 1 (ACAT-1).

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6.3. Mission Integration Process

6.3.1. Integration Meetings

The core of the mission integration process consists of a series of Mission Integration and Range
Working Groups (MIW G and RWG, respec tively). The MIWG has responsibility for all ph ysical interfaces
between the payload and t he launch vehicle. As suc h, the MIWG develops the Pa yload-to-Minotaur ICD
in addition to all miss ion-unique analyses, hardware, sof tware, and integrated procedures. The RW G is
responsible for items associated with launch site operations. Examples of such items include range
interfaces, hazardous pr ocedures, sys tem safe ty, and trajectory design. Docum entation produced by the
RWG includes all required range and safety submittals.

Working Group membership consists of the Mission Manager and representatives from Minotaur
engineering and operat ions organizations, as well as their counterparts from the c ustomer organization.
Quarterly meetings are typical, however the number of meetings re quired to develop and implem ent the
mission integration process will vary based on the complexity of the spacecraft.

6.3.2. Mission Design Reviews (MDR)

Two mission-specific design reviews will be held to determine the status and adequacy of the launch
vehicle mission prepar ations. They are designated MDR-1 an d MDR-2 and are typically held 6 months
and 13 months, res pectively, after authorit y to procee d. They are e ach analogo us to Prelim inary Design
Reviews (PDRs) and Cr itic al Design R evie ws (CD Rs), but focus prim aril y on mis sion-spec ific elem ents of
the launch vehicle effort.

6.3.3. Readiness Reviews

During the integration process, readiness reviews are held to provide the coordination of mission
participants and gain approval t o proceed to the next phase of ac tivity from senior m anagement. D ue to
the variability in complexity of different payloads, missions, and mission assurance categories, the
content and number of these reviews are tailored t o customer requirem ents. A brief descripti on of each
readiness review is provided below:

a. Pre-Ship Readiness Review (PSRR) — Conducted prior to committing flight hardware and

personnel to the f ield. T he PSR R provides testi ng res ults o n all f orm al s ystems tests and rev iews

the major mechanical ass em blies which are com plete d and r ead y for shipp ing at least L-60 da ys.

Safety status and field ope rations planning are also p rovided covering Range flight term ination,

ground hazards, spaceport coordination status, and facility preparation and readiness.
b. Incremental Readiness Review (IRR) – The quantity and timing of IRR(s) depends on the

complexity and Mission As surance Category of the mission. IRRs typically oc cur 2-12 months

prior to the launch date. IRR provides an early assessment of the integrated launch

vehicle/payload/facility readiness.
c. Mission Readiness Review (MRR) — Conduc ted within 2 months of launch, the MRR prov ides

a pre-launch assessment of integrated launch vehicle/payload/facility readiness prior to

committing significant resources to the launch campaign.
d. Flight Readiness Review (FRR) – The FRR is conducted at L-10 days and determines the

readiness of the integrated launch vehicle/payload/facility for a safe and successful launch. It

also ensures that all flight and ground hardware, software, personnel, and procedures are

operationally ready.
e. Launch R eadiness Review (LRR) — The LRR is conduct ed at L-1 day and serves as the final

assessment of mission readiness prior to activation of range resources on the day of launch.

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6.4. Documentation

Integration of the pa yload requires detailed, complete, and tim ely preparation and submittal of interface
documentation. SD/S DL is the primary comm unication path with other U.S. Gove rnment agencies , which
include—but are not limited to—the various Rang es and their support agenc ies, the U.S. Departm ent of
Transportation, U.S. State Department, and U.S. Department of Defense. The major products and
submittal times ass ociated with these organizations are di vided into two areas—those products that are
provided by the custom er, and those produce d by Orbital. Customer-provided d ocuments represent th e
formal communication of requirements, safety data, system descriptions, and mission operations
planning.

6.4.1. Customer-Provided Documentation

Documentation produced by the customer is detailed in the following paragraphs.

6.4.1.1. Payload Questionnaire

The Payload Questio nnaire is designe d to provide the initial definition of payload requirem ents, interface
details, launch site facilities, and preliminary safety data. Prior to the Mission Kickoff Meeting, the
customer shall provide the information requested in the Payload Questionnaire form (Appendix A).
Preliminary payload dra wings, as we ll as an y other pertine nt inform ation, sho ul d a l s o be inc luded with the
response. The customer’s responses to the payload questionnaire define the most current payload
requirements and interfaces and are instrumental in Orbital’s preparation of numerous documents
including the ICD, Prelim inary Mission Analyses and launch range docum entation. Orbital understands
that a definitive respons e to some questions m ay not be feasible prior to the Mission Kickoff Meeting as
they will be defined during the course of the mission integration process.

6.4.1.2. ICD Inputs

The LV-to-payload ICDs ( mission, mechanical and elec trical) detail all the mission s pecific requirements
agreed upon by Orbit al and the c ustomer. T hese key docum ents are used to ensure the c ompatibilit y of
all launch vehicle and payload interfaces, as well as defining all mission-specific and payload- unique
requirements. As such, the customer defines and provides to Orbital all the inputs that relate to the
payload. These inputs include those required to support flight trajectory development (e.g., orbit
requirements, pa yload mass pr operties, and payload s eparatio n requirem ents), mec hanical and e lectrica l
interface definition, payload unique requirements, payload operations, and ground support requirements.

6.4.1.3. Payload Mass Properties

Payload mass propert ies must be provided in a tim ely manner in order to support eff icient launch vehicle
trajectory developm ent and dynamic anal yses. Preliminary mass properties should be subm itted as part
of the MRD at launch vehicle authority to proceed. Updated mass properties shall be provided at
predefined intervals identified during the initial mission integration process. Typical timing of these
deliveries is included in Figure 6-2.

6.4.1.4. Payload Finite Element Model

A payload mathematical model is required for use in Orbital’s preliminary coupled loads analyses.
Acceptable form s include either a Craig-Bampton m odel valid to 120 Hz or a NAST RAN finite element
model. For the final coupled loads analysis, a test verified mathematical model is desired.

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6.4.1.5. Payload Thermal Model for Integrated Thermal Analysis

An integrated therm al analysis can be performed for any payload as a non-standard service. A pa yload
thermal model will be required from the payload organization for use in Orbital’s integrated thermal
analysis if it is required. The analysis is conducted for three mission phases:

a. Prelaunch ground operations;
b. Ascent from lift-off until fairing jettison; and
c. Fairing jettison through payload deployment.

The preferred thermal model format is Thermal Desktop, although FEMAP and SINDA/G can also be
provided. There is no limit on model size; however, larger models may increase the turn-around time.

6.4.1.6. Payload Drawings

Orbital prefers electronic versions of payload configuration drawings to be used in the mission specific
interface control dr awing, if poss ible. Orbital will work with the c ustom er to define the co ntent an d des ired
format for the drawings.

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

In order to obtain range support, a PRD must be prepared. This document describes requirements
needed to generally support the Minotaur launch vehicle. For each launch, an annex is submitted to
specify the range support needed to meet the mission’s requirem ents. This annex includes all payload
requirements as well as any additional Minotaur requirements that may arise to support a particular
mission. The customer completes all appropriate PRD forms for submittal to Orbital.

6.4.1.7.1. Launch Operations Requirements (OR) Inputs

To obtain range suppor t for the launch operation and as sociated rehearsals, an OR m ust be prepared.
The customer m ust pr ovide al l pa yload pr e-launc h a nd la unch da y requir em ents for inc orpor ation into the
mission OR.

6.4.1.8. Payload Launch Site Integration Procedures

For each mission, Orbital requires detailed spacecraft requirements for integrated launch vehicle and
payload integration acti vities. W ith these requirements, Or bital will produce the in tegrated procedures for
all launch site activiti es. In addition, al l payload procedur es that are perf ormed near the LV (e ither at the
integration facility or at the launch site or both) must be presented to Orbital for review prior to first use.

6.4.1.9. ICD Verification Documentation

Orbital conducts a rigor ous verification program to ensure all requ irements on both sides of the launch
vehicle-to-payload interface have been successfully fulfilled. As part of the ICD, Orbital includes a
verification matrix that indicates how each ICD requirement will be verified (e.g., test, analysis,
demonstration, etc.). As par t of the verification process, Orbital will pr ovide the customer with a matrix
containing all interface req uirements that are the responsibil ity of the payload to m eet. The matrix clearly
identifies the docum entation to be provided as proof of verif ication. Likewise, Orbital will e nsure that the
customer is provided with similar data for all interfaces that are the responsibility of launch vehicle to
verify.

6.4.2. Orbital Produced Documentation, Data, and Analyses

Mission documentation produced by Orbital is detailed in the following paragraphs.

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6.4.2.1. Launch Vehicle to Payload ICD

The launch vehicle-to-pa yload ICD detai ls all of the mission-uni que requirem ents agreed u pon by Orb ital
and the customer. The ICD is a critical docum ent used to ensure compatibilit y of all launch vehicle and
payload interfaces, as well as defining all mission-specific and mission-unique requirements. The ICD
contains the payload description, electrical and mechanical interfaces, environmental requirements,
targeting parameters, mission-peculiar vehicle requirement description, and unique GSE and facilities
required. As a critical part of this document, Orbital provides a comprehensive matrix that lists all ICD
requirements and the method in which these requirements are verified, as well as who is responsible.

The launch vehicle to payload ICD, as well as the Payload Mechanical ICD and Electrical ICD are
configuration controlled documents that are approved by Orbital and the customer. Once released,
changes to these docum ents are f ormally issue d and appr ove d by both part ies. The ICDs are reviewe d in
detail as part of the MIWG process.

6.4.2.2. ICD Verification Documentation

Orbital conducts a rigorou s verification program to ensure all requ irements on both sides of the launch
vehicle-to-payload interf ace have been successf ully fulfilled. Like the cus tomer-provided verif ication data
discussed in Section 6.4.1.9, Or bital will provide the c ustomer with similar data f or all interfaces that ar e
the responsibilit y of la unch vehicle to verif y. This documentation is us ed as part of the team eff ort to show
that a thorough verification of all ICD requirements has been completed.

6.4.2.3. Preliminary M ission Analyses

Orbital performs preliminary mission analyses to determine the compatibility of the payload with the
Minotaur launch vehicle and to pr ovide succinct, detailed m ission requirements such as lau nch vehicle
trajectory information, performance capability, accuracy estim ates and preliminary mission sequencing.
Much of the data derived from the preliminary mission analyses is used to establish the ICD and to
perform initial range coordination.

6.4.2.4. Coupled Loads Analyses (CLA)

Orbital has develop ed and validated fin ite element structural models of th e Minotaur vehicle f or use in
CLAs with Minotaur payloads. Orbital will incorporate the customer-provided payload model into the
Minotaur finite element model and perform a preliminary CLA to determine the maximum responses of the
entire integrated stack under transient loads. O nce a test validated spac ecraft model has been de livered
to Orbital, a f inal CL A load c ycle is com pleted. Through close coord inatio n betwe en the cus tom er and th e
Orbital, interim results can be made available to support the customer’s schedule critical needs.

6.4.2.5. Integrated Launch Site Procedures

For each mission, Orbit al prepares int egrated procedu res for various o perations that i nvolve the pa yload
at the processing facility and launch site. These include, but are not limited to: payload mate to the
Minotaur launch vehicle; f airing encapsulation; mission sim ulations; final vehic le closeouts, and transp ort
of the integrated launc h vehicle/payload to the launch pad. Once customer inputs are rec eived, Orbital
will develop draft proc e dur e s for review and comm ent. O nc e conc ur renc e is r eac h ed, f ina l procedures will
be released prior to use. Draft hazardous procedure s must be presented to the appr opriate launch site
safety organization 90 days prior to use and final hazardous procedures are due 45 days prior to use.

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6.4.2.6. Missile System Pre-Launch Safety Package (MSPSP) Annex

The MSPSP Annex doc um ents l aunch v ehicle and payload safety information inc lud ing an assessment of
any hazards which m ay arise from mission-specific vehicle and/or payload f unctions, and is pr ovided as
an annex to the basel ine Minotaur MSPSP. The c ustom er must pr ovide Orbita l with al l safet y inform ation
pertaining to the payload. Orbital assesses the c ombined vehicle and payload for hazards a nd pr epares a
report of the findings. Or bit al will t hen f or war d th e int e gr ated ass es s ment to the appropriate launc h Range
for approval.

6.4.2.7. PRD Mission Specific Annex

Once customer PRD input s are received, Orbital rev iews the inputs and upon resolving any concerns or
potential issues, submits the mission specific PRD annex to the range for approval. The range will
respond with a Program Support Plan (PSP) indicating their ability to support the stated requirements.

6.4.2.8. Launch Operation Requirements (OR)

Orbital submits the OR to obtain range support for pre-launch and launch operations. Information
regarding all aspects of launch day, particularly communication requirements, is detailed in the OR.
Orbital generates the document, solicits comments from the customer, and, upon comment resolution,
delivers the miss ion OR to the r ange. T he range ge nerates the Operat ions Direc tive (OD) th at is us ed by
range support personnel as the instructions for providing the pre-launch and launch day services.

6.4.2.9. Mission Constraints Document (MCD)

This Orbital-produced d ocument summarizes launch day operatio ns for the Minotaur launc h vehicle as
well as for the payload. Included in this document is a comprehensive definition of the Minotaur and
payload launch operati ons constraints, the established cr iteria for each constraint, the decision m aking
chain of command, and a summary of personnel, equipment, communications, and facilities that will
support the launch.

6.4.2.10. Final Countdown Procedure

Orbital produces the launc h countdown procedure that readi es the Minotaur launch vehic le and payload
for launch. All Minotaur and payload final countdown activities are included in the procedure.

6.4.2.11. Post-Laun ch Analyses

Orbital provides pos t-launc h anal yses to the c ustom er in two forms . The f irst is a quic k-look assessm ent
provided within four days of launch. The quick-look data report includes preliminary trajectory
performance data, orbital accuracy estimates, system performance preliminary evaluations, and a
preliminary assessment of mission success.

The second post-lau nch analysis, a m ore detailed f inal report of th e miss ion, is provided to the customer
within 30 days of launch. Included in the final mission report are the actual mission trajectory, event tim es ,
significant events, e nv iron ments, orbital param eter s a nd other pertinent d ata f ro m on-board telem etry and
Range tracking sensors. Photographic and video documentation, as available, is included as well.

Orbital also analyzes telemetry data from each launch to validate Minotaur performance against the
mission ICD requirem ents. In the case of any miss ion anomaly, Orbital will cond uct an investigation and
closeout review.

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6.5. Safety

6.5.1. System Safety Requirements

In the initial phases of the m ission integration effort, regulat ions and instruc tions that appl y to spacecraf t
design and processing are reviewed. Not all safety regulations will apply to a particular mission
integration activit y. Tailoring the range requirem ents to the mis sion unique activities will be the first step
in establishing the safety plan.

Before a spacecraft arr ives at the processing site, the payloa d organization must provide the cogn izant
range safety offic e with certification that the system has been designed and test ed in accordance with
applicable safety requirements (e.g. AFSPCM 91-710 for CCAFS and VAFB). Spacecraft must also
comply with the sp ecific pa yload proc essing f acilit y s afet y requirem ents. Orbital will prov ide the c ustomer
assistance and guidance regarding applicable safety requirements.

It cannot be overstressed that the applicable safety requirements should be considered in the earliest
stages of spacecraft design. Processing and launch site ranges discourage the use of waivers and
variances. Furthermore, approval of such waivers cannot be guaranteed.

6.5.2. System Safety Documentation

For each Minotaur mission, Orbital acts as the interface with Range Safety. In order to fulfill this role,
Orbital requires safety information from the payload. For launches from either the Eastern or Western
Ranges, AFSPCM 91-710 provides detailed range safety regulations. To obtain approval to use the
launch site facilities, specific data m us t be prepared and s ubm itted to Orbital. This inform ation includes a
description of each payload ha zardous system and evidence of c ompliance with safety requ irements for
each system. Drawings, s chem atics , and ass em bly and handl ing proc edur es, incl uding pro of tes t data for
all lifting equipm ent, as well as any other inform ation that will aid in assessing the respective systems
should be included. Major categories of ha zardous systems ar e ordnance devices, radioact ive materials,
propellants, pressuri zed systems, toxic materials, cr yogenics, and RF radiation. Procedures rel ating to
these systems as well as any procedures relating to lifting operations or battery operations should be
prepared for safet y review submittal. Orbital will provide this inform ation to the appropriate saf ety offices
for approval.

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Figure 7-1. Hardware Flow – Factory to Launch Site

7. GROUND AND LAUNCH OPERATIONS

Minotaur ground and launch operations processing minimizes the handling complexity for both launch
vehicle and payload. A ll launch vehicle m otors, parts and com pleted subassemblies ar e delivered to the
Minotaur Processing Fac ility (MPF) from either Orbital’s Chandler pr oduction facilit y, the assembly/m otor
vendor, or the Government. Ground and launch operations are conducted in three major phases:

a. Launch Vehicle Integration — Assembly and test of the Minotaur lau nc h veh icle .
b. Payload Processing/Integration — Receipt and checkout of the payload, followed by integration

with the Minotaur launch vehicle interface, verification of those interfaces and payload

encapsulation.
c. Launch Operations — Inc ludes transport to the launch pad, f inal integration, check out, arming

and launch.

Figure 7-1 depicts the typical flow of hardware from the factory to the launch site.

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Integration at MPF

7.1. Launch Vehicle Integration Overview

Orbital utilizes the same f undamental integration an d process f low for all launch vehicles in the Minotaur
family. A flowchart of the launch vehicle integration at the MPF is shown in Figure 7.1-1 for a VAFB
Minotaur IV launch. The f low to accommodate other Minotaur config urations or other launch facilities is
similar and modified as required. The timeline described in this section pertains to a nominal launch
campaign. Figure 7.1-2 sho ws the Minotaur hardware and support equ ipment undergoing integrat ion at
the MPF.

Figure 7.1-1. Launch Vehicle Processing Flow at the MPF

7.1.1. Planning and Documentation

Minotaur integration and test activities are c ontrolled by a compr ehensive set of Work Packages (WPs)
that describe and docum ent every aspect of integrati ng and testing the Minotaur launch v ehicle and its
payload. All testing and integration activities are scheduled by work package number on an activity
schedule that is updated and distributed daily during field operations. This schedule is maintained by
Orbital and serves as the master document
communicating all activities planned at the field
site. The schedule contains notations regarding
the status of the work package document and
hardware required to begin the operation.
Mission-specific work packages are created for
mission-unique or payload-specific procedures.
Any discrepancies enc ountered are r ecorded on a
Non-Conformance Report and dispositioned as
required. All activities are in accordance with
Orbital’s ISO 9001 certification.

7.1.2. Guidance and Control Assembly Integration and Test Activities

The Guidance and Control Assembly (GCA) will
undergo system level testing at Orbita l’s Chandler
facility prior to being s hipped to the field si te. The

Figure 7.1-2. Minotaur Launch Vehicle

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GCA and the upper sta ge motor are then del ivered to the MPF located at VAFB. Upon arrival at VAFB
these assemblies wil l undergo a thorough inspection and subs ystem level checkout. At this time range
certification of Range Tracking System (RTS) and Flight Termination System (FTS) devices will be
performed at both the component and in-vehicle testing level. After the com pletion of subsystem level
testing, the motor is integrated into the GCA to form the GCA/motor assembly.

7.1.3. PK Motor Integration and Test Activities

The PK motors ar e delivered to the MPF where they undergo checkout, integration, and testing. These
activities include ordnance and raceway installation, as well as steering and phasing tests.

7.1.4. Mission Simulation Tests

Orbital will r un at least two Mission Simulation Tests (MST) to verify the functionality of launch vehicle
hardware, and software. T he Miss ion Sim ulation T ests use the actual f light sof twar e and sim ulate a “f ly to
orbit” scenario using simulated Inertial Navigation System (INS) data. This allows the test to proceed
throughout all mission phases and capture vehicle performance data. The data will be compared to
previous MSTs perform ed in the factory using the same flight software and har dware. Orbital developed
PK Thrust Vector Actuator (TVA) simulators are used to perform all mission simulations. These
components provide a realistic as sessment of booster perf ormance dur ing the te sting operati ons. After a
thorough data review of all telemet ry parameter s, the tes t configurati on is disass embled and prep ared for
payload integration.

7.1.5. Launch Vehicle Processing Facilities

The Minotaur Processing Facility (MPF), Building
1900, at VAFB is a 48,000 sq. ft facility used
primarily for LV processing prior to transporting the
LV to the appropri ate launch site or range f or that
mission. For mis sions out of VAFB, the MPF has
adequate floor space an d infrastructure to support
concurrent launch vehicle and pa yload processin g.
The MPF is shown in Figure 7.1.5-1. Should the
MPF be utilized for payload processing, it is
expected that the payload and Minotaur launch
vehicle would be processed in separate sec ti ons of
the High Bay area.

The MPF has infrastructure capability to s upport payload processing requirem ents in terms of security,
electrical and communications service, overhead crane, and a temperature and humidity controlled
environment. High Cleanliness operations are discussed further in 8.2.3.1 as required per the mission
and particle containment requirements.

Figure 7.1.5-1. Minotaur Processing Is

Performed at the MPF at VAFB

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7.2. Payload Processing/Integration

Payloads typically u ndergo initial checkout and preparation f or launch at a Payload Processing Facility
(PPF), which can be either government provided or commercial facility. After arrival at the PPF (see
Figure 7.2-1), the payload completes its own independent verification and checkout prior to beginning
integrated processing with the Minotaur launch vehicle. When integrated processing is ready to
commence, the Minotaur fairing and Payload Adapter Module (PAM) are delivered to the payload
processing facilit y. T he p a yload is mounted to the PAM and th en encapsu lated b y the fairing, as shown i n
Figure 7.2-2. The encapsul ated assembly is then s hipped in the vert ical configuration t o the launch site,
as shown in Figure 7.2-3, where it will underg o pre-stack verification test. Toget her, the f airing an d PAM
provide a sealed assembly which protects the payload during transport and launch.

Figure 7.2-1. Payload Processing and LV Integration Flow at the PPF

Figure 7.2-2. Payload Encapsulation at the PPF

Figure 7.2-3. Encapsulated Payload Transport

to the Launch Site

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7.2.1. Payload Propellant Loading

Payloads utilizing integral propulsion systems with propellants such as hydrazine can be loaded and
secured through coordinated OSP arrangements. This is a non-standard service.

7.3. Launch Operations

At the completion of activities at the MPF and PPF, the final phase of the Launch c ampaign is entered.
This begins with the stack ing of the booster stages and culm inates with the launch of the Minotaur a nd
payload. A notional lau nch operations f low chart is shown in F igure 7.3-1. The L-m inus dates may var y
from mission to mission depending on vehicle configuration and other range commitments. Launch
operations activities are described in more detail in the subsections to follow.

Figure 7.3-1. Minotaur IV Launch Site Operations

7.3.1. Booster Assembly Stacking/Launch Site Preparation

After completion of the launch vehicle testing at the MPF, the booster stages and the GCA/motor
assembly are transported to the launch facility.

Prior to the arrival of the P K bo os ters , the site is prepared for la unc h o per at io ns with t he ins ta llati on of t he
launch stand adapter.

Each PK motor is individ ually transported down to the launc h site. Once a motor arrives at the launch
site, it is rolled off the transporter and then rotated into a vertical configuration. It is then lifted and
emplaced onto the launch stand adapter. This process is repeated for each PK stage.

The GCA/motor assembly is shipp ed in the vertical co nfiguration to the launc h site, where it is emplaced
on top of the PK motor s tack. Stacking operati ons are shown in Figure 7.3.1-1 as performed at V AFB
SLC-8 for a Minotaur IV mission.

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Figure 7.3.1-1. Minotaur Uses Vertical Integration for Each Booster Stage, the Guidance Control

Assembly, and the Encapsulated Payload Assembly

7.3.2. Final Vehicle Integration and Test

After successful completion of payload mate and fairing closeout, the encapsulated payload is
transported to the p ad in a vertical configurati on and then lifted atop the booster assembly (see Figure

7.3.1-1). Final post-mate checks of the booster assembly and front section assembly interface are
conducted, followed b y a final system s verification tes t. At this point th e vehicle is ready for final R ange
interface tests.

7.3.3. Launch Vehicle Arming

Following final vehic le testi ng, the lau nch veh icle is ar m ed and the pad is cleared f or launch. T he majority
of these arming activities occur at L-1 day and bring the Minotaur launch vehicle nearly to its launch day
configuration. L-1 da y is als o typica lly the last opport unity for payload ac cess. The l ast remaini ng arming
steps (final arming) occur mid-way during the countdown on launch day.

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Figure 7.3.4-2. Minotaur IV Prepared for Launch

7.3.4. Launch

The typical Minotaur final countdown procedure commenc es at 5 hours prior to the required launch time.
Figure 7.3.4-1 describes the n ominal Minotaur launch day flow. T hese activities methodically trans ition
the vehicle from a safe s tate to that of launch r eadines s. Pa yload task s , as necessar y, are includ ed in the
countdown procedur e and are coord inated b y the M inotaur Lau nch Cond uctor. T he Minota ur IV is sho wn
ready for launch in Figure 7.3.4-2.

Figure 7.3.4-1. Notional Minotaur Countdown Timeline

7.3.5. Launch Control Organization

The Launch Control Orga nization is split into two
groups: the Managem ent group and the T echnica l
group. The Management gr oup consists of senior
range personnel and Mission Director s/Managers
for the launch vehicle and payload who provide
authority to proceed at selected points in the
countdown. The Technical Group consists of the
Launch Vehicle, Payload and Range personnel
responsible for executio n of the launch operation,
to include data review and launch readiness
assessment. The Payload’s members of the
technical group are engineers who provide
technical represent ation in the control c enter. The
Launch Vehicle’s m embers of the technical group
are engineers who pre pare the Minotaur for flight,
review and assess data that is displayed in the
Launch Control Room (LCR) and provide
technical representation in the LCR and in the
Launch Operations Control Center (LOCC). The
Range’s members of the technical group are
personnel that maintai n and m onitor the v oice a nd
data equipment, tracking facilities and all assets
involved with RF comm unications with the launch
vehicle. I n addition, the Ra nge provides personnel
responsible for the Flight Termination System
monitoring and commanding.

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7.3.6. Launch Rehearsals

Two rehearsals are con ducted prior to each launch. The fir st is conducted at approximately L-10 days
and is used to acquaint th e launch team with the communications s ystems, reporting, problem solving,
launch procedures and constraints, and the decision making process. The first rehearsal is
communications on ly ( i. e., the Minotaur l aunch vehicle and pa yload are not powered on and range as set s
are not active). It is typically a full day in durati on and consists of a number of countdowns perf ormed
using abbreviated timelines, clock jumps, and off-nominal situations. All aspects of the team’s
performance are exer cised, as well as hold, scr ub, and r ecycle proc edures. T he opera tions are c ritiqued
and the lessons learned are incorporated prior to the Mission Dress Rehearsal (MDR) at L-5 days
(typical). The MDR is the final rehearsal prior to the actual launch day operation. It will ensure that
problems encountered during the first rehearsal have been resolved. The MDR exercises the entire 5
hour Minotaur count down procedure an d simulated post launch events. The Launch Vehicle is powered
for this rehearsal and range assets perform operations as they would on launch day. There are no
planned off-nominal e vents, ho wever the team will rea ct to real world anom alies as the y would on launch
day. MDR ends with successful completion of the countdown procedure.

All Customer personnel involved with launch day activities participate in both rehearsals.

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Additional Access Panel

8. OPTIONAL ENHANCED CAPABILITIES

The OSP launch service is structured to provide a baseline vehicle configuration which is then augmented
with optional enhancements to meet the unique needs of individual payloads. The baseline vehicle
capabilities are define d in t he pre vio us sec tions an d th e opti on al e nha nc ed capa bi lit ies ar e def in ed b elo w.
The enhanced options allow customization of launch support and accommodations to the Minotaur
designs on an efficient, “as needed” basis.

8.1. Separation System and Optional Mechanical Interfaces

Several different types of opti onal separation systems and mechanic al interfaces are avai lable through
Orbital. Further details can be found in Sections 5.2.4 and 5.2.5.

8.2. Conditioned Air

Conditioned air is included in the baseline vehicle
cost and was described previously in Section 4.6.1.
The Nitrogen Purge and Enhanced Contamination
Control enhancements complement this capability
as described in the enhanc ements Section 8.3 and

8.6.

8.3. Nitrogen Purge

Clean, dry gaseous nitrogen (GN
Grade B specificatio ns as defined in MIL-P-27401C
can be provided to the payload in a Class 10,000
environment for continuous purge of the payload
after fairing encapsulation until final payload
closeouts (non-fly away) or until lift-off (flyaway
configuration shown in Figure 8.3-1). This
enhancement uses a flow regulated nitr ogen grou nd
supply connected to the fairing. The nitrogen flow
control regulator ensur es the purge is supplied at a
minimum flow rate of 5 standard cubic feet per
minute with a capability of up to 8 standard cubic
feet per minute. A manifold mounted t o the inside of
the fairing wall feeds lines up the fairing wall to
purge points of interest on the payload. Purge
nozzles can be positioned on the fairing wall and
pointed at the payload i ns tr ument. Alternatively, a fly
away configuration can be used where the purge
line connects to a manifold on the payload and is
pulled free during fairing separation. This continuous
purge can be supplied from payload encapsulation
through launch, including during transport to the
pad.

8.4. Additional Access Panel

As already discussed in Section 5.1.3, additional
doors of the same size and configuration as the
standard single access door can be provided. The

) purge meeting

2

Figure 8.3-1. GN

Purge Interface To Minotaur

2

Fairing (Flyaway at Liftoff)

Figure 8.4-1. Example Location and Size of

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Only)

location of the fairing ac cess door is document ed within the mission-s pecific ICD. The allowabl e access
door envelopes are sho wn in Figure 5.1.3-1. Required door locations outside the all owable envelope as
shown in Figure 8.4-1 are evaluated on a miss ion­specific basis. Other f airing acces s configuratio ns,
such as small circular access panels, can be
provided as non-standard, mission-specific
enhancements. Additional mission-specific effort
can be minimized if a previously flown access
door configuration is chosen.

8.5. Enhanced Telemetry

Enhanced telemetr y provides for m ission specific
instrumentation and telemetry components to
support additional payload, LV, or experiment
data acquisition requirem ents. This enhancement
provides a dedicated telem etry link with a baseline
data rate of 2 Mbps. Additi onal instrumentation or
signals such as strain gauges, temperature
sensors, accelerom eters, analog\ and digital data
can be configured to meet mission specific
requirements. This capability was successfully
demonstrated on the first five Minotaur IV
launches. Typical enhanced telemetry
instrumentation includes accelerometers (ECA)
and microphones (ECM) intended to capt ure high
frequency transients such as shock and random
vibration. A sample of the enhanced telemetry
instrumentation location on the Minotaur 92”
payload fairing is provided in Figure 8.5-1.

8.6. Enhanced Contamination Control

To meet the requirement f or a low contamination
environment, Orbital uses existing processes
developed and demonstrated on the Minotaur,
Taurus, and Pegasus progr ams. Thes e process es
are designed to minimize out-gassing, supply a
Class 10,000 clean room environment, assure a
high cleanliness pa yload envelope, and provide a
HEPA-filtered, controlled humidity environment
after fairing encapsulation. Orbital leverages
extensive payload processing experience to
provide flexible, responsive solutions to mission­specific payload requirements (Figure 8.6-1).

8.6.1. Low Outgassing Materials

Orbital’s existing high cleanliness design and
integration processes ensure that all materials

Figure 8.5-1. Representative Minotaur 92”

Enhanced Instrumentation Locations (Fairing

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used within the encapsulated volume have
outgassing character istics of less than 1.0% Total
Mass Loss (TML) and less than 0.1% Collected
Volatile Condensable Mass (CVCM) in
accordance with ASTM E59. If m ater ials withi n the
encapsulated volum e cannot m eet low outgas sing
characteristics because of unique mission
requirements, a contamination control plan is
developed to ensure controls are in place to
eliminate any significant effect on the payload.

8.6.2. High Cleanliness Integration Environment

With the enhanced contamination control option,
the encapsulated payload element of the vehicle
is processed in an ISO Standard 14644-1 Class
10,000 environm ent during all p ayload proces sing
activities up to fairing encapsulation (ISO 7). The
Payload Processing Facility (PPF) clean room
(Figure 1.6.6-2) utilizes HEPA filtration units to
filter the air and ensure hydrocarbon content is
maintained at ≤15 ppm, with humidity maintained
at 30-60% relative humidity. Depending on
payload requirements , the c lean room can als o be
certified as Class 100,000 (ISO 8) while still
providing tighter environmental control than the
standard high-bay environment, thereby streamlining access and payload processing.

8.6.3. HEPA-Filtered Fairing Air Supply

With the enhanced contamination control option, the ECU continuously purges the fairing volume with
clean filtered air while maintaining temperatur e, humidity, and cleanl iness. Orbital’s ECU incor porates a
HEPA filtration unit a long w ith a h ydroc arbo n filter adaptor to provide C lass 10 ,00 0 ( ISO 7) a ir and ensure
hydrocarbon content is maintained at ≤15 ppm, with humidity maintained as stated in section 4.6.1.
Orbital monitors the suppl y air f or partic ulate matter via a probe inst alled ups tream of the fairin g inlet duct
prior to connecting the air source to the payload fairing.

8.6.4. Fairing Surface Cleanliness

The inner surfac e of the fai ring and exp osed launch v ehicle assem blies are cleaned t o Visibly Clean Plus
Ultraviolet cleanliness criteria which ensures no particulate matter visible with normal vision when
inspected from 6 to 18 inches under 100 foot candle incident light, as well as when the surface is
illuminated by black light at 3200 to 3800 Angstroms. Process and procedures for inspection and the
bagging of material to preclude contamination during shipment to the field are in place.

8.7. Secure FTS

The Secure FTS (Figure 8.7-1) is achieved with the L-3 Cincinnati Electronics Model CRD-120/205
Launch Vehicle Command Receiver/Decoder that is compatible with the "High-Alphabet" range safety
modulation format. The receiver uses a pre-stored code unique to each specific vehicle to issue

Figure 8.6-1. Minotaur Team Has Extensive

Experience in a Payload Processing Clean

Room Environment

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Figure 8.7-1. Orbital’s Secure FTS System Block Diagram

configuration and term ination commands. This provid es an increased level of securit y over the standard
FTS systems that use a basic 4 tone combination for receiver command and control.

The CRD-120/205 Launch Veh icle Command Receive r/Decoder was designed s pecifically to operate on
the Delta expendable spac e launch vehicles for range saf ety flight termination. This design inc orporates
redundancy in bot h hardware and sof tware and H igh Reliabilit y piece-parts (in accordance w ith ELV-JC­002D) to ensure reliable, fail-safe operation.

8.8. Over Horizon Telemetry

A Telemetry Data R el a y Sa t ellit e S ystem (TDRSS) interface can be added as an enhancement to pr o v ide
real-time telemetr y cover ag e duri ng b lac k out peri ods with ground based telem etry receiving sites. TDRSS
was successfull y demonstrated on past Mi notaur missions. The TDRSS telemetr y system enhancement
consists of a LCT2 T DRSS transm itter , an antenna (Figur e 8.8-1) , one RF switch, and as sociat ed gr ound
test equipment. The RF switch is used dur ing ground tes ting to allo w for a test ant enna to be used in l ieu
of the flight antennas . Near the tim e when telem etry covera ge is lost b y ground b ased tel em etr y receiving
sites, the LV switches telemetry output to the
TDRSS antenna and poi nts the antenna to wards a
TDRSS satellite. The T DRSS relays the telemetry
to the ground where it is then route d to the launch
control room (Figure 8.8-2). A cavity backed or
phased array antenna can be used depending on
data rate requirements. The TDRSS system
proposed includes the launch vehicle design,
analysis, hardware a nd launch vehicle testing. For
this option, arrangements need to be made with
NASA for system support and planning,

Figure 8.8-1. TDRSS 20W LCT2 Transmitter and

Cavity Backed S-band Antenna

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