ORBITAL SCIENCES CORPORATION

January 2006 Minotaur IV User’s Guide

Release 1.1

Approved for Public Release Distribution Unlimited

Copyright 2006 by Orbital Sciences Corporation. All Rights Reserved.

Minotaur IV User’s Guide Revision Summary

REVISION SUMMARY

VERSION DOCUMENT DATE CHANGE PAGE 1.0 TM-17589 Jan 2005 Initial Release All

1.1 TM-17589A Jan 2006 General nomenclature, history, and administrative updates All (no technical updates) • Updated launch history • Corrected contact information

Release 1.1 January 2006 ii Minotaur IV User’s Guide Preface

This User’s Guide is intended to familiarize payload mission planners with the capabilities of the Orbital Suborbital Program ll (OSP-2) Minotaur IV Space Launch Vehicle (SLV) launch service. The information 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) for the payloader organization to use in their development/design process. This document provides an overview of the Minotaur IV system design and a description of the services provided to our customers.

USAF SMC Det 12/RP 3548 Aberdeen Ave SE Kirtland AFB, NM 87117-5778

Telephone: (505) 846-8957 Fax: (505) 846-5152

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

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

Telephone: (480) 814-6566 E-mail: [email protected]

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Release 1.1 January 2006 iv Minotaur IV User’s Guide Table of Contents

PAGE 1. INTRODUCTION...... 1 2. MINOTAUR IV LAUNCH SERVICE ...... 3 2.1. Minotaur IV Launch System Overview...... 3 2.2. Minotaur IV Launch Service...... 3 2.3. Minotaur IV Launch Vehicle...... 4 2.3.1. Stage 1, 2 and 3 Booster Assemblies...... 4 2.3.2. Stage 4 Booster/Avionics Assembly ...... 5 2.3.2.1. Avionics...... 5 2.3.2.2. Attitude Control Systems ...... 5 2.3.2.3. Telemetry Subsystem ...... 7 2.3.2.4. Payload Fairing ...... 7 2.4. Launch Support Equipment ...... 8 2.4.1. Transportable LSE Shelters...... 8 3. GENERAL PERFORMANCE ...... 11 3.1. Mission Profiles...... 11 3.2. Launch Sites ...... 11 3.2.1. Western Launch Sites ...... 11 3.2.2. Eastern Launch Sites ...... 11 3.2.3. Alternate Launch Sites ...... 11 3.3. Performance Capability...... 11 3.4. Injection Accuracy...... 23 3.5. Payload Deployment...... 23 3.6. Payload Separation...... 23 3.7. Collision/Contamination Avoidance Maneuver ...... 23 4. PAYLOAD ENVIRONMENT...... 25 4.1. Steady State and Transient Acceleration Loads...... 25 4.1.1. Transient Loads...... 26 4.1.2. Steady-State Acceleration...... 26 4.2. Payload Vibration Environment ...... 26 4.2.1. Random Vibration ...... 28 4.2.2. Sine Vibration...... 28 4.3. Payload Acoustic Environment ...... 28 4.4. Payload Shock Environment...... 28 4.5. Payload Structural Integrity and Environments Verification...... 32 4.5.1. Recommended Payload Testing and Analysis ...... 33 4.6. Thermal and Humidity Environments...... 33

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PAGE 4.6.1. Ground Operations...... 33 4.6.2. Powered Flight ...... 34 4.6.3. Nitrogen Purge (Non-Standard Service) ...... 35 4.7. Payload Contamination Control ...... 35 4.8. Payload Electromagnetic Environment...... 35 5. PAYLOAD INTERFACES...... 37 5.1. Payload Fairing ...... 37 5.1.1. Payload Dynamic Design Envelope...... 37 5.1.2. Payload Access Door...... 37 5.2. Payload Mechanical Interface and Separation System ...... 37 5.2.1. Standard Non-Separating Mechanical Interface...... 38 5.2.2. Orbital Supplied Mechanical Interface Control Drawing ...... 38 5.3. Payload Electrical Interfaces...... 38 5.3.1. Payload Umbilical Interfaces...... 38 5.3.2. Payload Interface Circuitry ...... 40 5.3.3. Payload Battery Charging ...... 40 5.3.4. Payload Command and Control...... 40 5.3.5. Pyrotechnic Initiation Signals ...... 40 5.3.6. Payload Telemetry ...... 40 5.3.7. Payload Separation Monitor Loopbacks ...... 41 5.3.8. Telemetry Interfaces ...... 41 5.3.9. Non-Standard Electrical Interfaces ...... 41 5.3.10. Electrical Launch Support Equipment...... 41 5.4. Payload Design Constraints...... 41 5.4.1. Payload Center of Mass Constraints ...... 41 5.4.2. Final Mass Properties Accuracy...... 41 5.4.3. Pre-Launch Electrical Constraints...... 42 5.4.4. Payload EMI/EMC Constraints...... 42 5.4.5. Payload Dynamic Frequencies ...... 42 5.4.6. Payload Propellant Slosh...... 42 5.4.7. Payload-Supplied Separation Systems...... 42 5.4.8. System Safety Constraints...... 42 6. MISSION INTEGRATION...... 43 6.1. Mission Management Approach ...... 43 6.1.1. RSLP Mission Responsibilities...... 43 6.1.2. Orbital Mission Responsibilities ...... 43

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PAGE 6.2. Mission Planning and Development ...... 44 6.3. Mission Integration Process...... 44 6.3.1. Integration Meetings...... 44 6.3.2. Mission Design Reviews (MDR)...... 45 6.3.3. Readiness Reviews...... 45 6.4. Documentation...... 46 6.4.1. Customer-Provided Documentation...... 46 6.4.1.1. Payload Questionnaire...... 46 6.4.1.2. Payload Mass Properties...... 46 6.4.1.3. Payload Finite Element Model ...... 46 6.4.1.4. Payload Thermal Model for Integrated Thermal Analysis...... 46 6.4.1.5. Payload Drawings ...... 46 6.4.1.6. Program Requirements Document (PRD) Mission Specific Annex Inputs ...... 46 6.4.1.6.1. Launch Operations Requirements (OR) Inputs ...... 47 6.5. Safety ...... 47 6.5.1. System Safety Requirements...... 47 6.5.2. System Safety Documentation...... 47 7. GROUND AND LAUNCH OPERATIONS ...... 49 7.1. Minotaur IV/Payload Integration Overview ...... 49 7.2. Ground And Launch Operations ...... 49 7.2.1. Launch Vehicle Integration...... 49 7.2.1.1. Planning and Documentation...... 49 7.2.1.2. GCA/Orion 38 Integration and Test Activities...... 49 7.2.1.3. PK Motor Integration and Test Activities...... 49 7.2.1.4. Mission Simulation Tests ...... 49 7.2.1.5. Booster Assembly Stacking/Launch Pad Preparation ...... 51 7.2.2. Payload Processing/Integration ...... 51 7.2.2.1. Payload Propellant Loading...... 52 7.2.2.2. Final Vehicle Integration and Test ...... 52 7.3. Launch Operations...... 53 7.3.1. Launch Control Organization ...... 53 8. OPTIONAL ENHANCED CAPABILITIES...... 55 8.1. Mechanical Interface and Separation System Enhancements...... 55 8.1.1. Separation Systems ...... 55 8.1.2. Additional Fairing Access Doors ...... 55 8.1.3. Payload Isolation System...... 55

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PAGE 8.2. Performance Enhancements ...... 56 8.2.1. Insertion Accuracy...... 56 8.2.2. Star 48 Stage 4 ...... 57 8.3. Environmental Control Options...... 58 8.3.1. Conditioned Air...... 58 8.3.2. Nitrogen Purge ...... 58 8.3.3. Enhanced Contamination Control ...... 58 8.3.3.1. High Cleanliness Integration Environment (Class 10K or 100K)...... 59 8.3.3.2. Fairing Surface Cleanliness Options ...... 59 8.3.3.3. High Cleanliness Fairing Environment...... 59 8.3.4. Launch Pad Environmental Control ...... 59 8.3.4.1. Booster Temperature Control ...... 59 8.4. Enhanced Telemetry Options ...... 60 8.4.1. Enhanced Telemetry Bandwidth ...... 60 8.4.2. Enhanced Telemetry Instrumentation ...... 60 8.4.3. Navigation Data...... 60 8.5. Shared Launch Accommodations...... 60

LIST OF FIGURES Figure 2-1. OSP-2 Peacekeeper Space Lift Vehicle...... 3 Figure 2-2. Minotaur IV Expanded View ...... 4 Figure 2-3. Orion 38 Stage 4 Motor ...... 6 Figure 2-4. Stage 4 Structures...... 6 Figure 2-5. Existing 92 in. Taurus Fairing, Handling Fixtures, and Processes will be used for Minotaur IV...... 7 Figure 2-6. Functional Block Diagram of LSE...... 9 Figure 2-7. Portable Launch Support Structure Provide Optional Support From Austere Sites...... 10 Figure 3-1. Minotaur IV Mission Profile...... 12 Figure 3-2. Minotaur IV Launch Site Options...... 13 Figure 3-3. Minotaur IV Performance Curves for VAFB Launches (Metric Units) ...... 14 Figure 3-4. Minotaur IV Performance Curves for VAFB Launches (English Units) ...... 15 Figure 3-5. Minotaur IV Performance Curves for CCAFS Launches (Metric Units) ...... 16 Figure 3-6. Minotaur IV Performance Curves for CCAFS Launches (English Units) ...... 17 Figure 3-7. Minotaur IV Performance Curves for Kodiak, Alaska Launches (Metric Units)...... 18 Figure 3-8. Minotaur IV Performance Curves for Kodiak, Alaska Launches (English Units)...... 19 Figure 3-9. Minotaur IV Performance Curves for Wallops, Virginia Launches (Metric Units)...... 20 Figure 3-10. Minotaur IV Performance Curves for Wallops, Virginia Launches (English Units)...... 21

Release 1.1 January 2006 viii Minotaur IV User’s Guide Table of Contents

PAGE Figure 3-11. Stage Impact Points for VAFB and CCAFS Launches...... 22 Figure 3-12. Minotaur IV Injection Accuracy...... 24 Figure 3-13. Typical Pre-Separation Payload Pointing and Spin Rate Accuracies ...... 24 Figure 4-1. Phasing of Dynamic Loading Events...... 25 Figure 4-2. Payload CG Net Transient Lateral Acceleration at Stage 2 Ignition with a Typical Separation System...... 27 Figure 4-3. Minotaur IV Nominal Maximum Axial Acceleration as a Function of Payload Mass ...... 27 Figure 4-4. Minotaur IV Payload Random Vibration Environment...... 29 Figure 4-5. Minotaur IV Payload Sine Vibration Environment ...... 29 Figure 4-6. Minotaur IV Payload Acoustic Maximum Predicted Environment (MPE)...... 30 Figure 4-7. Minotaur IV Payload Shock Maximum Predicted Environment (MPE) – Launch Vehicle to Payload ...... 31 Figure 4-8. Payload Shock Environment – Payload to Launch Vehicle ...... 32 Figure 4-9. Factors of Safety Payload Design and Test ...... 33 Figure 4-10. Recommended Payload Testing Requirements...... 33 Figure 4-11. Payload Thermal and Humidity Environment ...... 34 Figure 4-12. Minotaur IV Launch Vehicle RF Emitters and Receivers ...... 36 Figure 5-1. Standard 92 in. Fairing Envelope ...... 37 Figure 5-2. Standard, Non-separating Payload Mechanical Interface...... 39 Figure 5-3. Payload Electrical Interface Block Diagram, With No Orbital Supplied Separation System .. 39 Figure 5-4. Payload 1:1 Umbilical Pin Outs ...... 39 Figure 5-5. Minotaur IV Payload Electrical Interface Block Diagram...... 40 Figure 5-6. Payload Mass Properties Measurement Tolerance ...... 42 Figure 6-1. Typical Integrated OSP Organizational Structure ...... 43 Figure 6-2. Typical Mission Integration Schedule...... 45 Figure 7-1. Hardware Flow – Factory to Launch Site ...... 50 Figure 7-2. SLV Processing Flow ...... 51 Figure 7-3. Minotaur IV Processing Flow...... 52 Figure 7-4. Minotaur IV Upper Stack Assembly will be Vertically Integrated to Minotaur IV Booster Assembly in a Similar Manner to Taurus Upper Stack ...... 53 Figure 8-1. 38-in. Separation System Option...... 56 Figure 8-2. Soft Ride Payload Isolation System as Integrated on Minotaur LV...... 57 Figure 8-3. Hydrazine Auxiliary Propulsion System (HAPS) Used to Provide Insertion Accuracy...... 57 Figure 8-4. Orion 38 Stage 4 Motor can be Replaced with a Star-48 to Provide Increased Performance 58 Figure 8-5. Mobile Scaffolding for Environmental Control Demonstrated on Minotaur Missions ...... 59 Figure 8-6. Modular Minotaur IV Structural Design Easily Accommodates Multiple Payloads ...... 60

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LIST OF TABLES PAGE TABLE 4-1. PAYLOAD CG PARAMETRIC DESIGN LIMIT LOADS...... 26

LIST OF APPENDICES A. PAYLOAD QUESTIONNAIRE...... A-1

Release 1.1 January 2006 x Minotaur IV User’s Guide Glossary

ACS Attitude Control System MACH Modular Avionics Control Hardware A/D Arm/Disarm MARS Mid-Atlantic Regional Spaceport AADC Alaska Aerospace Development MDR Mission Design Review Corporation MIWG Mission Integration Working Groups AC Air Conditioning MRD Mission Requirements Document AFRL Air Force Research Laboratory MRR Mission Readiness Review BER Bit Error Rate NTW Navy Theater Wide C/CAM Collision/Contamination Avoidance ODM Ordnance Driver Modules Maneuver OR Operations Requirements CCAFS Cape Canaveral Air Force Station OSP-2 Orbital Suborbital Program ll CDR Critical Design Review PAF Payload Attach Fitting CG center-of-gravity PAM Payload Adapter Module CLA Coupled Loads Analysis PBCM PK Booster Control Module CLF Commercial Launch Facility PCF Programmable Clock Filter CMP Critical Measurements Program PCM Pulse Code Modulation DPAF Dual Payload Attach Fitting PDR Preliminary Design Review ECS Environmental Control System PID Proportional-Integral-Derivative EELV Evolved Expendable Launch Vehicle PK Peacekeeper EGSE Electrical Ground Support Equipment POC Point of Contact EMC Electromagnetic Compatibility PPF Payload Processing Facility EME Electromagnetic Environment PRD Program Requirements Document EMI Electromagnetic Interference PWB Printed Wiring Board FLSA Florida Spaceport Authority QES Quick Erect Scaffold FTS Flight Termination System RF Radio Frequency GCA Guidance and Control Assembly RSLP Rocket Systems Launch Program GFE Government Furnished Equipment RTS Range Tracking System GFP Government Furnished Property RWG Range Working Groups GPB GPS Position Beacon SEA Statistical Energy Analysis GPS Global Positioning System SEB Support Equipment Building GTO Geosynchronous Transfer Orbit SES Saab Ericson Space HAPS Hydrazine Auxiliary Propulsion SLV Space Launch Vehicle System SRSS Softride for Small Satellites HEPA High Efficiency Particulate Air SSI Spaceport Systems International HVAC Heating, Ventilation and Air START Strategic Arms Reduction Treaty Conditioning TPS Thermal Protection System ICD Interface Control Drawing TVAs Thrust Vector Actuators LCR Launch Control Room TVC Thrust Vector Control LER Launch Equipment Room USAF United States Air Force LEV Launch Equipment Van VAFB Vandenberg Air Force Base LOCC Launch Operations Control Center VPF Vehicle Processing Facility LRR Launch Readiness Review WFF Wallops Flight Facility LSE Launch Support Equipment WP Work Package LSV Launch Support Van LV Launch Vehicle MAC Motor Adapter Cone MAC Modified Adapter Cone

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Release 1.1 January 2006 xii Minotaur IV User’s Guide Section 1.0 - Introduction

1. INTRODUCTION This User’s Guide is intended to familiarize By adopting the austere launch site concepts payload mission planners with the capabilities of developed for Taurus, the Minotaur IV system can the Orbital Suborbital Program ll (OSP-2) Minotaur operate from a wide range of launch facilities and IV Space Launch Vehicle (SLV) launch service. geographic locations. The system is compatible The information provided in this user’s guide is for with, and will typically operate from, commercial initial planning purposes only. Information for spaceport facilities and existing U.S. Government development/design is determined through ranges at Vandenberg Air Force Base (VAFB), mission specific engineering analyses. The Cape Canaveral Air Force Station (CCAFS), results of these analyses are documented in a Wallops Flight Facility (WFF), and Kodiak Island, mission-specific Interface Control Document (ICD) Alaska. This User’s Guide describes Minotaur IV- for the payloader organization to use in their unique integration and test approaches (including development/design process. This document the typical operational timeline for payload provides an overview of the Minotaur IV system integration with the Minotaur IV vehicle) and the design and a description of the services provided ground support equipment that will be used to to our customers. Minotaur IV offers a variety of conduct Minotaur IV operations. enhanced options to allow the maximum flexibility in satisfying the objectives of single or multiple payloads.

The primary mission of the Minotaur IV is to provide low cost, high reliability launch services to government-sponsored payloads. Minotaur IV accomplishes this using flight-proven components with a significant flight heritage such as surplus Peacekeeper boosters, the Taurus Fairing and Attitude Control System, and a mix of Minotaur I, Pegasus, Taurus, and other orbital standard avionics, all with a proven, successful track record. The philosophy of placing mission success as the highest priority is reflected in the success and accuracy of all OSP missions to date.

The Minotaur IV launch vehicle system is composed of a flight vehicle and ground support equipment. Each element of the Minotaur IV system has been developed to simplify the mission design and payload integration process and to provide safe, reliable space launch services. This User’s Guide describes the basic elements 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 environments, and outlines the Minotaur IV mission integration process.

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Release 1.1 January 2006 2 Minotaur IV User’s Guide Section 2.0 – Minotaur IV Launch Service

2. MINOTAUR IV LAUNCH SERVICE IV draws on the successful heritage of four launch vehicles: Orbital’s Pegasus, Taurus and Minotaur I 2.1. Minotaur IV Launch System Overview space launch vehicles and the Peacekeeper The Minotaur IV (Figure 2-1) mission is to system of the USAF. Minotaur IV’s avionics are provide a cost effective, reliable and flexible derived from the Pegasus and Taurus systems, means of placing satellites into orbit. The launch providing a combined total of more than 35 vehicle developer and manufacturer is Orbital, successful space launch missions. Orbital’s under the Orbital Suborbital Program 2 (OSP-2) efforts have enhanced or updated Pegasus, contract for the U.S. Air Force. An overview of the Taurus and Minotaur avionics components to meet system and available launch services is provided the payload-support requirements of the OSP-2 within this section, with specific elements covered program. Combining these improved subsystems in greater detail in the subsequent sections of this with the long successful history of the User’s Guide. 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 complete 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 portable and adaptable to varying levels of infrastructure. While the Minotaur IV system is capable of self-contained operation using portable vans to house the EGSE, it is typically launched from an established range where the EGSE can be housed in available, permanent structures or facilities.

The vehicle and LSE are designed to be capable of launch from any of the four commercial Spaceports (Alaska, California, Florida, and Mid-Atlantic), as well as from existing U.S. Government facilities at VAFB and CCAFS. The Launch Control Room (LCR) serves as the control Figure 2-1. OSP-2 Peacekeeper Space Lift center for conducting a Minotaur IV launch and Vehicle includes consoles for Orbital, range safety, and limited customer personnel. Further description of Minotaur IV has been designed to meet the the Launch Support Equipment is provided in needs of U.S. Government-sponsored customers Section 2.4. at a lower cost than commercially available alternatives by the use of surplus Peacekeeper 2.2. Minotaur IV Launch Service boosters. The requirements of the OSP-2 The Minotaur IV Launch Service is provided program stress system reliability, transportability, through the combined efforts of the USAF and and operation from multiple launch sites. Minotaur Orbital, along with associate contractors including

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Northrop Grumman and Commercial Spaceports. The primary customer interface will be with the USAF Space and Missile Systems Center, Detachment 12, Rocket Systems Launch Program, designated RSLP. Orbital is the launch vehicle provider. For brevity, this integrated team effort will be referred to as “OSP”. Where interfaces are directed toward a particular member of the team, they will be referred to directly (i.e., Orbital or RSLP).

OSP provides all of the necessary hardware, software and services to integrate, test and launch a satellite into its prescribed orbit. In addition, OSP will complete all the required agreements, 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. This provides a solid basis for initiating and streamlining the integration process for future Minotaur IV customers.

2.3. Minotaur IV Launch Vehicle The Minotaur IV vehicle, shown in expanded view in Figure 2-2, 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 simplified integration and test capability, yield a robust, highly reliable launch vehicle design. In addition, Minotaur IV payload accommodations and interfaces have been 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 consists of the refurbished Government Furnished Equipment (GFE) Peacekeeper Stages 1, 2 and 3. These booster assemblies are used as provided by the Government, requiring no modification or Figure 2-2. Minotaur IV Expanded View additional components. They have extensive flight history, including 51 Peacekeeper launches and propellant motors and utilize a movable nozzle three Stage 1 launches on Taurus with no motor controlled by a Thrust Vector Actuator (TVA) related failures. All three stages are solid- system for three-axis attitude control. The first

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stage consists of a Thiokol motor that provides flight-proven Modular Avionics Control Hardware 500,000 lbf (2224 kN) of thrust. The second stage (MACH). Standardized, function-specific modules motor is an Aerojet motor with a moveable nozzle are combined in stacks to meet vehicle-specific contoured with an extendable exit cone. It requirements. The functional modules from which provides an average thrust of 275,000 lbf the MACH stacks are created include power (1223 kN). The third stage is a Hercules motor transfer, ordnance initiation, booster interface, that provides 65,000 lbf (289 kN) of thrust and also communication, and telemetry processing. Orbital features an extendable exit cone similar to has designed, tested, and flown a variety of MACH Stage 2. modules, which provide an array of functional capability and flexibility. MACH has exhibited 2.3.2. Stage 4 Booster/Avionics Assembly 100% reliability on all flights including Minotaur The Minotaur IV Stage 4 motor is the Orion 38 and TLV flights and several of Orbital’s suborbital design (Figure 2-3). This motor was originally launch vehicles including Navy Theater Wide developed for Orbital’s Pegasus program and is (NTW), Critical Measurements Program (CMP), used on the Minotaur I launch vehicle, as well as and STORM. other Orbital launch vehicles. Common design features, materials and production techniques are 2.3.2.2. Attitude Control Systems applied to the motors to maximize reliability and The Minotaur IV Attitude Control System production efficiency. The Orion 38 motor (ACS) provides attitude control throughout provides the velocity needed for orbit insertion for boosted flight and coast phases. Stages 1, 2 and 3 the SLV, in the same manner as it is used on the utilize the PK Thrust Vector Control (TVC) Minotaur. This motor features state-of-the-art systems. The PK Booster Control Module (PBCM) design and materials with a successful flight links the flight computer actuator commands to the heritage. It is currently in production and is individual Stage 1, 2, and 3 Thrust Vector actively flying payloads into space, with 37 fully Actuators (TVAs). Stage 4 utilizes the same TVC successful flights to date and one static test. system used by the Pegasus, Taurus and Minotaur vehicles which combines single-nozzle Three substructures are utilized to electromechanical TVC for pitch and yaw control accommodate the Orion 38 Stage 4 motor and with a three-axis, cold-gas attitude control system attach it to the Stage 1-3 PK booster assembly. resident in the avionics section providing roll These are a Payload Adapter Module (PAM) with control. 62.01-inch diameter payload interface, a 38-inch diameter motor adapter cone and a GCA +3/4 Attitude control is achieved using a three-axis interstage. These structures were adapted from autopilot that employs Proportional-Integral- similar Taurus hardware designs and are shown in Derivative (PID) control. Stages 1, 2 and 3 fly a Figure 2-4. pre-programmed attitude profile based on trajectory design and optimization. Stage 4 uses a 2.3.2.1. Avionics set of pre-programmed orbital parameters to place The Minotaur IV avionics system has heritage the vehicle on a trajectory toward the intended to the Minotaur I, OSP TLV, as well as Pegasus insertion apse. The extended coast between and Taurus designs. The flight computer is a 32- Stages 3 and 4 is used to orient the vehicle to the bit multiprocessor architecture. It provides appropriate attitude for Stage 4 ignition based communication with vehicle subsystems, the LSE, upon a set of pre-programmed orbital parameters and the payload, if required, utilizing standard RS- and the measured performance of the first three 422 serial links and discrete I/O. The avionics stages. Stage 4 utilizes energy management to system design incorporates Orbital’s innovative, place the vehicle into the proper orbit. After the

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Figure 2-3. Orion 38 Stage 4 Motor

Figure 2-4. Stage 4 Structures

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final boost phase, the three-axis cold-gas attitude control during ground handling, integration control system is used to orient the vehicle for operations and flight. The fairing is a bi-conic spacecraft separation, contamination and collision design made of graphite/epoxy face sheets with avoidance and downrange downlink maneuvers. an aluminum honeycomb core. The fairing The autopilot design is modular, so additional provides for low payload contamination through payload requirements such as rate control or prudent design and selection of low contamination celestial pointing can be accommodated with materials and processes. Acoustic blankets and minimal additional development. internal air conditioning ducts are available to provide more benign payload environments. Air 2.3.2.3. Telemetry Subsystem conditioning will keep the payload environment to The Minotaur IV telemetry subsystem provides a specified temperature between 60 to 120 °F real-time health and status data of the vehicle dependent upon requirements. avionics system, as well as key information regarding the position, performance and The two halves of the fairing are structurally environment of the Minotaur IV vehicle. This data joined along their longitudinal interface using may be used by Orbital and the range safety Orbital’s low contamination frangible joint system. personnel to evaluate system performance. The An additional circumferential frangible joint at the Minotaur IV baseline telemetry subsystem base of the fairing supports the fairing loads. At provides a number of dedicated payload discrete separation, a gas pressurization system is (bi-level) and analog telemetry monitors through activated to pressurize the fairing deployment dedicated channels in the SLV encoder. The thrusters. The fairing halves then rotate about Minotaur IV telemetry system provides a baseline external hinges that control the fairing deployment 1 Mbps data rate (both payload and Minotaur IV to ensure that payload and launch vehicle telemetry). However, the output data rate is clearances are maintained. All elements of the selectable from 2.441 kbps to 10 Mbps to allow deployment system have been demonstrated flexibility in supporting evolving mission through test to comply with stringent requirements, as limited by link margin and Bit contamination requirements. Error Rate (BER) requirements. The telemetry subsystem nominally utilizes Pulse Code Options for payload access doors and Modulation (PCM) with a RNRZ-L format. enhanced cleanliness are available. Further However other types of data formats, including NRZ-L, S, M, and Bi-phase may be implemented if required, in order to accommodate launch range limitations.

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

2.3.2.4. Payload Fairing Orbital’s flight-proven Taurus 92-inch diameter Figure 2-5. Existing 92 in. Taurus Fairing, payload fairing (Figure 2-5) is used to encapsulate Handling Fixtures, and Processes the payload, provide protection and contamination will be used for Minotaur IV

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details on the baseline fairing are included in payload consoles and equipment can be Section 5.1. With the addition of a structural supported in the LCR and SEB, within the adapter, the fairing can accommodate multiple constraints of the launch site facilities or temporary payloads. This feature, described in more detail in structure facilities. Interface to the payload Section 8.5 of this User’s Guide, permits two or through the Minotaur IV payload umbilicals and more smaller payloads to share the cost of a landlines provides the capability for direct Minotaur launch, resulting in a lower launch cost monitoring of payload functions. Payload for each as compared to other launch options. personnel accommodations will be handled on a OSP has access to several Multiple Payload mission-specific basis. Adaptor (MPA) designs that allow for a cost sharing benefit to programs with excess payload 2.4.1. Transportable LSE Shelters and/or mass capability. In order to perform mission operations from alternative, austere launch sites, Orbital can 2.4. Launch Support Equipment provide self contained, transportable shelters for The Minotaur IV LSE is designed to be readily the Launch Support Van (LSV) and Launch adaptable to varying launch site configurations Equipment Van (LEV) as an unpriced option. with minimal unique infrastructure required. The These shelters are the same approach and design EGSE consists of readily transportable consoles used on all six of Orbital’s Taurus launches that can be housed in various facility (Figure 2-7). The OSP-2 Ground Support configurations depending on the launch site Consoles have been intentionally made modular infrastructure. The EGSE is composed of three and portable to allow their use in these primary functional elements: Launch Control, accommodations. Vehicle Interface, and Telemetry Data Reduction. The Launch Control consoles are located in a The LSV consists of a shelter which is located Launch Control Room (LCR), or mobile launch at a Range Safety-approved man-safe distance equipment van depending on available launch site from the launch site. The LSV contains the vehicle accommodations. The Vehicle Interface EGSE is control and telemetry monitor consoles described located in a permanent structure, typically called a in Section 2.4. Sufficient space is available for Support Equipment Building (SEB) or Launch additional equipment racks depending on Equipment Room (LER). Fiber optic connections Government and/or Payload requirements. The from the Launch Control to the Vehicle Interface LEV consists of an 8 foot x 20 foot shelter which is consoles are used for efficient, high bandwidth located near the launch stool and is unmanned communications and eliminates the need for during launch. This shelter contains the vehicle copper wire. The Vehicle Interface racks provide interface racks described in Section 2.4. The LEV the junction from fiber optic cables to the copper has sufficient room available for payload power cabling interfacing with the vehicle. Figure 2-6 supplies and interface electronics. Both shelters depicts the functional block diagram of the LSE. are designed for shipping and transportation with exterior tiedown and anchor locations used to The LCR serves as the control center during facilitate the loading and unloading operations. the launch countdown. The number of personnel The shelters can be delivered to any level location that can be accommodated are dependent on the and be set up within hours. The transportable launch site facilities. At a minimum, the LCR will support console design also allows for the LSE to accommodate Orbital personnel controlling the be moved into a fixed blockhouse and LER if vehicle, two Range Safety representatives (ground required. and flight safety), and the Air Force Mission Manager. Mission-unique, customer-supplied

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Figure 2-6. Functional Block Diagram of LSE

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Figure 2-7. Portable Launch Support Structure Provide Optional Support From Austere Sites

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3. GENERAL PERFORMANCE 3.2.2. Eastern Launch Sites For Easterly launch azimuths to achieve 3.1. Mission Profiles orbital inclinations between 28.5° and 60°, Minotaur IV can attain a range of posigrade Minotaur IV can be launched from facilities at and retrograde inclinations through the choice of Cape Canaveral, FL or Wallops Island, VA. launch sites made available by the readily Launches from Florida will nominally use the adaptable nature of the Minotaur IV launch Florida Spaceport Authority (FLSA) launch system. A typical mission profile to a sun- facilities at LC-46 on CCAFS. These will be synchronous orbit is shown in Figure 3-1. All typically for inclinations from 28.5° to 40°, although performance parameters presented herein are inclinations above 35° may have reduced typical for most expected payloads. However, performance due to the need for a trajectory performance may vary depending on unique dogleg. Inclinations below 28.5° are feasible, payload or mission characteristics. Specific albeit with doglegs and altitude constraints due to requirements for a particular mission should be range safety considerations. The Mid-Atlantic coordinated with OSP. Once a mission is formally Regional Spaceport (MARS) facilities at the WFF initiated, the requirements will be documented in may be used for inclinations from 30° to 60°. the Mission Requirements Document (MRD). Some inclinations may have reduced performance Further detail will be captured in the Payload-to- due to range safety considerations and will need Launch Vehicle Interface Control Drawing (ICD). to be evaluated on a case-by-case mission- specific basis. 3.2. Launch Sites Depending on the specific mission and range 3.2.3. Alternate Launch Sites safety requirements, Minotaur IV can operate from Other launch facilities can be readily used several East and West Coast launch sites, given the flexibility designed into the Minotaur IV illustrated in Figure 3-2. Specific performance vehicle, ground support equipment, and the parameters are presented in Section 3.3. various interfaces. Orbital has experience Facilities used for OSP Minotaur and Taurus launching vehicles from a variety of sites around launches are generally compatible with Minotaur the world. To meet the requirements of IV operations. performing mission operations from alternative, austere launch sites, Orbital can provide self 3.2.1. Western Launch Sites contained, transportable shelters as described in For missions requiring high inclination orbits section 2.4.1. (greater than 60°), launches can be conducted from facilities at VAFB or Kodiak Island, AK. Both 3.3. Performance Capability facilities can accommodate inclinations from 60° to Minotaur IV performance curves for circular 120°, although inclinations below 72° from VAFB and elliptical orbits of various altitudes and require an out-of-plane dogleg, thereby reducing inclinations are detailed in Figure 3-3 through payload capability. As with the initial Minotaur Figure 3-10 for launches from all four Spaceports missions, Minotaur IV can be launched from the in both metric and English units. These California Spaceport facility, Space Launch performance curves provide the total mass above Complex 8 (SLC-8) operated by Spaceport the standard, non-separating interface. The mass Systems International (SSI), on South VAFB. The of any Payload Attach Fitting (PAF) or separation launch facility at Kodiak Island, operated by the system is to be accounted for in the payload mass Alaska Aerospace Development Corporation allocation. Figure 3-11 illustrates the stage (AADC) has been used for both orbital and vacuum impact points for launch trajectories from suborbital launches. VAFB and CCAFS.

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Figure 3-1. Minotaur IV Mission Profile

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Figure 3-2. Minotaur IV Launch Site Options

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Figure 3-3. Minotaur IV Performance Curves for VAFB Launches (Metric Units)

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Figure 3-4. Minotaur IV Performance Curves for VAFB Launches (English Units)

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Figure 3-5. Minotaur IV Performance Curves for CCAFS Launches (Metric Units)

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Figure 3-6. Minotaur IV Performance Curves for CCAFS Launches (English Units)

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Figure 3-7. Minotaur IV Performance Curves for Kodiak, Alaska Launches (Metric Units)

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Figure 3-8. Minotaur IV Performance Curves for Kodiak, Alaska Launches (English Units)

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Figure 3-9. Minotaur IV Performance Curves for Wallops, Virginia Launches (Metric Units)

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Figure 3-10. Minotaur IV Performance Curves for Wallops, Virginia Launches (English Units)

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Figure 3-11. Stage Impact Points for VAFB and CCAFS Launches

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3.4. Injection Accuracy Separation velocities are driven by the need to Minotaur IV injection accuracy is summarized prevent recontact between the payload and the in Figure 3-12. Better accuracy can be provided Minotaur IV upper stage after separation. The dependent on specific mission characteristics. For value will typically be 2 to 3 ft/sec (0.6 to 0.9 example, heavier payloads will typically have m/sec). better insertion accuracy, as will higher orbits. An enhanced option for increased insertion accuracy 3.7. Collision/Contamination Avoidance is also available (Section 8.2.1). It utilizes the Maneuver flight-proven Hydrazine Auxiliary Propulsion Following orbit insertion and payload System (HAPS) developed on the Pegasus separation, the Minotaur IV Stage 4 will perform a program. Collision/Contamination Avoidance Maneuver (C/CAM). The C/CAM minimizes both payload 3.5. Payload Deployment contamination and the potential for recontact Following orbit insertion, the Minotaur IV between Minotaur IV hardware and the separated Stage 4 avionics subsystem can execute a series payload. OSP will perform a recontact analysis for of ACS maneuvers to provide the desired initial post-separation events. payload attitude prior to separation. This capability may also be used to incrementally A typical C/CAM begins soon after payload reorient Stage 4 for the deployment of multiple separation. The launch vehicle performs a 90° spacecraft with independent attitude requirements. yaw maneuver designed to direct any remaining Either an inertially-fixed or spin-stabilized attitude Stage 4 motor impulse in a direction which will may be specified by the customer. The maximum increase the separation distance between the two spin rate for a specific mission depends upon the bodies. After a delay to allow the distance spin axis moment of inertia of the payload and the between the spacecraft and Stage 4 to increase to amount of ACS propellant needed for other a safe level, the launch vehicle begins a “crab- attitude maneuvers. Figure 3-13 provides the walk” maneuver to impart a small amount of delta typical payload pointing and spin rate accuracies. velocity, increasing the separation between the payload and the fourth stage of the Minotaur IV. 3.6. Payload Separation Payload separation dynamics are highly Following the completion of the C/CAM dependent on the mass properties of the payload maneuver as described above and any remaining and the particular separation system utilized. The maneuvers, such as downlinking of delayed primary parameters to be considered are payload telemetry data (per START treaty provisions), the tip-off and the overall separation velocity. ACS valves are opened and the remaining ACS nitrogen propellant is expelled. 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. If an optional Orbital-supplied Marmon Clamp-band separation system is used, payload tip-off rates are generally under 5°/sec per axis. Separation system options are discussed further in Section 8.1.1. Orbital performs a mission-specific tip-off analysis for each payload.

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Tolerance Error Type Error Source (Worst Case) Altitude Stage 4 motor performance uncertainty and guidance ±10 nmi (18.5 km) (Insertion Apse) algorithm uncertainty Altitude Stage 4 motor performance and guidance algorithm ±50 nmi (92.6 km) (Non-Insertion Apse) uncertainty and navigation (INS) error Altitude Stage 4 motor performance and guidance algorithm ±30 nmi (55.6 km) (Mean) uncertainty and navigation (INS) error Guidance algorithm uncertainty and navigation Inclination ±0.2° (INS) error Figure 3-12. Minotaur IV Injection Accuracy

Figure 3-13. Typical Pre-Separation Payload Pointing and Spin Rate Accuracies

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4. PAYLOAD ENVIRONMENT events such as motor ignition, stage separation, and transonic crossover. CAUTION The environmental design and test criteria The predicted environments provided in this user's guide are for initial presented have been derived using measured planning purposes only. data obtained from previous Taurus, Minotaur I, and Peacekeeper missions, motor static fire tests, Environments presented here bound other system development tests and analyses. typical mission parameters, but should The predicted levels presented are intended to be not be used in lieu of mission-specific analyses. Mission-specific levels are representative of a standard mission. Satellite provided as a standard service and mass, geometry and structural components vary documented or referenced in the greatly and will result in significant differences mission ICD. from mission to mission.

This section provides details of the predicted Dynamic loading events that occur throughout environmental conditions that the payload will various portions of the flight include steady-state experience during Minotaur IV ground operations, acceleration, transient low frequency acceleration, powered flight, and launch system on-orbit acoustic impingement, random vibration, and operations. The predicted environments provided pyrotechnic shock events. Figure 4-1 identifies in this user’s guide are for initial planning purposes the time phasing of these dynamic loading events only. and environments and their significance. Pyroshock events are not indicated in this figure, Minotaur IV ground operations include payload as they do not occur simultaneous with any other integration and encapsulation within the fairing, significant dynamic loading events. In addition, subsequent transportation to the launch site and dynamic loading associated with S4 ignition is final vehicle integration activities. Powered flight insignificant. begins at Stage 1 ignition and ends at Stage 4 burnout. Minotaur IV on-orbit operations begin 4.1. Steady State and Transient Acceleration after Stage 4 burnout and end following payload Loads separation. To more accurately define Design limit load factors due to the combined simultaneous loading and environmental effects of steady state and low frequency transient conditions, the powered flight portion of the accelerations are largely governed by payload mission is further subdivided into smaller time characteristics. A mission-specific Coupled Loads segments bounded by critical, transient flight Analysis (CLA) will be performed, with customer

Figure 4-1. Phasing of Dynamic Loading Events

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provided finite element models of the payload, in vehicle model. A Monte-Carlo analysis is order to provide precise load predictions. Results performed to determine variations in vehicle will be referenced in the mission specific ICD. For acceleration due to changes in winds, motor preliminary design purposes, Orbital can provide performance and aerodynamics. The steady-state initial Center-of-Gravity (CG) netloads given a accelerations must be added to transient payload’s mass properties, CG location and accelerations from the CLA as indicated in bending frequencies. Design limit loads due to Figure 4-1 to determine the total payload both transient and steady-state accelerations are acceleration. Steady-state accelerations are presented in Table 4-1 for select payload masses. typically 8-Gs axial and 0.5-Gs lateral.

4.1.1. Transient Loads During powered flight, the maximum steady Transient loads account for approximately state accelerations are dependent on the payload 30% of the total vehicle load with the remainder mass. The maximum level occurs during Stage 3 due to steady wind loads. Typical acceleration burn. Figure 4-3 depicts the axial acceleration at values at the payload interface are 2-Gs lateral burnout for Stages 3 and 4 as a function of and 4-Gs axial depending on the load case. payload mass. Transient lateral accelerations at Stage 2 ignition are defined as a function of payload mass in 4.2. Payload Vibration Environment Figure 4-2. Preliminary and final CLAs will be The Minotaur IV payload vibration performed for each Minotaur IV payload. The environments are low frequency random and payload finite element model is coupled to the sinusoidal vibrations created by the solid rocket vehicle model. Forcing functions have been motor combustion processes and transmitted developed for different flight events or load cases. through the launch vehicle structure. Additionally, Load cases include liftoff, transonic, max q and higher frequency aeroacoustics energy is created stage and shroud separation. by air flow over the surface of the vehicle. Some of this aeroacoustic energy is transmitted via the 4.1.2. Steady-State Acceleration launch vehicle structure to the payload. However, The steady-state vehicle accelerations are a majority of the aeroacoustic energy is determined from the vehicle rigid body analysis. transmitted to the payload directly as acoustic Drag, wind and motor thrust are applied to a energy through the fairing.

TABLE 4-1. PAYLOAD CG PARAMETRIC DESIGN LIMIT LOADS

Payload Mass 1600 lbm (725.7 kg) 2400 lb (1089 kg) 3200 lb (1452 kg) 4000lb (8141 kg) Axial (G) Lateral Axial (G) Lateral Axial (G) Lateral Axial (G) Lateral (G) (G) (G) (G) max/min max/min max/min max/min Liftoff 3.83/0.27 0.62 3.93/0.15 0.46 3.90/0.16 0.41 4.01/0.12 0.37 Pre-Transonic Resonant Burn 5.05/0.83 0.02 3.46/2.29 0.00 3.44/2.31 0.00 3.73/2.08 0.00 Transonic 5.13/1.52 1.23 3.95/2.71 0.97 3.89/2.75 0.90 4.07/2.49 0.89 Supersonic 3.41/3.40 1.96 3.38/3.38 1.61 3.36/3.36 1.40 3.34/3.34 1.26 Stage 2 Ignition 3.93/-0.35 4.05 3.95/-0.03 2.89 3.83/-0.02 2.74 3.66/0.01 2.13 Stage 3 Ignition 6.79/0.00 0.78 6.45/0.00 0.59 6.22/0.00 0.49 5.90/0.00 0.40 Stage 3 Burnout See Figure 4-3 TBS See Figure 4-3 TBS See Figure 4-3 TBS See Figure 4-3 TBS Stage 4 Burnout See Figure 4-3 TBS See Figure 4-3 TBS See Figure 4-3 TBS See Figure 4-3 TBS

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Figure 4-2. Payload CG Net Transient Lateral Acceleration at Stage 2 Ignition with a Typical Separation System

Figure 4-3. Minotaur IV Nominal Maximum Axial Acceleration as a Function of Payload Mass

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4.2.1. Random Vibration Random vibration environments are produced reviewed test data from more than 30 static fire both by the structurally transmitted vibrations tests and three Taurus flights. The sine vibration through mechanical payload interface as well as level depends on payload weight and stiffness. from the acoustic energy directly through the Orbital has developed forcing function to simulate fairing (Figure 4-4). However, the higher frequency the Stage 1 motor resonance. The forcing aeroacoustic-induced random vibration levels are functions are used in a NASTRAN simulation to hard to accurately replicate for spacecraft of the predict peak sine vibration levels at different Minotaur IV class. Testing input levels are typically vehicle locations. The simulation has been hard to accurately specify because the response validated using flight data. The sine vibration at the LV-to-spacecraft interface is strongly varies between 45 and 75 Hz. The NASTRAN dependent on the unique spacecraft dynamics, simulation of the PK Stage 1 resonant burn is including its response to the acoustic field. analyzed for each payload as part of the CLA Therefore, structure-born random vibration discussed in Section 4.1. The results of this environments are only defined herein for mission-specific CLA are then used to refine the frequencies up to 250 Hz with the SLV vehicle and payload sine vibration levels. A recommendation that high frequency parametric analysis of hypothetical payloads has environments be based on the acoustic levels been conducted to provide guidance on expected defined in Section 4.3. The acoustic environment acceleration levels. The resulting levels are shown is defined starting at 20 Hz, allowing both in Figure 4-5. environments to be evaluated in the overlapping region. 4.3. Payload Acoustic Environment The acoustic environments to which the The structure-borne low frequency random spacecraft will be exposed have been defined vibration environment for Minotaur IV is a flat based on measured acoustic data from previous 0.002 g2/Hz from 20 to 250 Hz, producing an Taurus flights which utilized the Peacekeeper overall level of 0.68 gRMS (Figure 4-4). This level Stage 1 motor and same 92 in. fairing. The data is based on measured data from multiple was adjusted to account for differences in vehicle applicable Orbital launch vehicles. Vibration trajectories and a 3 dB uncertainty was added per produced by the PK Stage 1 motor was analyzed MIL-STD-1540 guidelines. The resulting acoustic based on data from three Taurus flights, which level is shown in Figure 4-6. utilized the PK Stage 1 motor. The random vibration levels also cover the Orion 38 motor burn 4.4. Payload Shock Environment observed on past Minotaur, Pegasus, and Taurus The maximum shock response spectrum at flights. Detailed evaluation of the Stage 2 and 3 the base of the payload from all launch vehicle motor burn is still pending, but prior experience events should not exceed the flight limit levels in indicates that their levels will be relatively benign Figure 4-7 (clamp-band separation system). for the payload. The Minotaur IV structure-borne Lower separation shock levels can be achieved by vibration environment is enveloped by the MIL- the use of a Lightband separation system provided STD-1540E levels. OSP recommends that the by Planetary Systems Incorporated. The resulting 1540E level be used as guidance for minimum levels are also shown in Figure 4-7. levels for payload testing. For missions that do not utilize an Orbital- 4.2.2. Sine Vibration supplied payload separation system, the shock A known resonant burn characteristic of the response spectrum at the base of the payload PK Stage 1 motor creates a sine vibration from vehicle events should not exceed the levels requirement of Minotaur IV payloads. Orbital has in Figure 4-7 (non-separating shock).

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Figure 4-4. Minotaur IV Payload Random Vibration Environment

Figure 4-5. Minotaur IV Payload Sine Vibration Environment

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Figure 4-6. Minotaur IV Payload Acoustic Maximum Predicted Environment (MPE)

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Figure 4-7. Minotaur IV Payload Shock Maximum Predicted Environment (MPE) – Launch Vehicle to Payload

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If the payload employs a non-Orbital a. Design Limit Load — The maximum separation system, then the shock delivered from predicted ground-based, powered flight or the separation systems to the Stage 4 vehicle on-orbit load, including all uncertainties. interface must not exceed the limit level b. Design Yield Load — The Design Limit characterized in Figure 4-8. Shock above this Load multiplied by the recommended Yield level could require requalification of components Factor of Safety (YFS). The payload or an acceptance of risk by the Rocket Systems structure must have sufficient strength to Launch Program (RSLP). withstand simultaneously the yield loads, applied temperature, and other 4.5. Payload Structural Integrity and accompanying environmental phenomena Environments Verification for each design condition without The primary support structure for the experiencing detrimental yielding or spacecraft must possess sufficient strength, permanent deformation. rigidity, and other characteristics required to c. Design Ultimate Load — The Design survive the critical loading conditions that exist Limit Load multiplied by the recommended within the envelope of handling and mission Ultimate Factor of Safety (UFS). The requirements, including worst-case predicted payload structure must have sufficient ground, flight, and post-boost loads. It must strength to withstand simultaneously the survive those conditions in a manner that assures ultimate loads, applied temperature, and safety and that does not reduce the mission other accompanying environmental success probability. Spacecraft design loads are phenomena without experiencing any defined as follows: fracture or other failure mode of the structure.

Figure 4-8. Payload Shock Environment – Payload to Launch Vehicle

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4.5.1. Recommended Payload Testing and Analysis Sufficient payload testing and/or analysis must be performed to ensure the safety of ground crews and to ensure mission success. The payload structural design should comply with the testing and design factors of safety in Figure 4-9. Vibration testing should be based on the standard margins defined in Figure 4-10. At a minimum, it is recommended that the following tests be performed: a. Structural Integrity — Static loads, sine vibration, or other tests should be performed that combine to encompass the acceleration load environment presented in Section 4.1. b. Random Vibration — The flight level environment and recommended test level is defined in Section 4.2.1. c. Acoustics — Full scale acoustic testing is recommended to verify higher frequency Figure 4-9. Factors of Safety Payload dynamics of the spacecraft are not Design and Test adversely affected. The acoustic levels are defined in Figure 4-6. d. Shock – The payload separation event should be simulated to verify the spacecraft is not adversely effected. Shock levels are defined in Section 4.4.

The payload organization must provide OSP with a list of the tests and test levels to which the Figure 4-10. Recommended Payload Testing payload was subjected prior to payload arrival at Requirements the integration facility. The HVAC provides conditioned air to the payload 4.6. Thermal and Humidity Environments in the PPF after fairing integration. The HVAC is The thermal and humidity environment to used at the launch pad after vehicle stacking which the payload may be exposed during vehicle operations. Air Conditioning (AC) is not provided processing and pad operations are defined in the during transport or lifting operations without the sections that follow and listed in Figure 4-11. enhanced option that includes High Efficiency Particulate Air (HEPA) filtration. The conditioned 4.6.1. Ground Operations air enters the fairing at a location forward of the Upon encapsulation within the fairing and for payload, exits aft of the payload and is provided the remainder of ground operations, the payload up to 5 minutes prior to launch. Baffles are environment will be maintained by a Heating, provided at the air conditioning inlet to reduce Ventilation and Air Conditioning (HVAC) system. impingement velocities on the payload if required.

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Figure 4-11. Payload Thermal and Humidity Environment

Fairing inlet conditions are selected by the than 0.1. This temperature limit envelopes the customer, and are bounded as follows: maximum temperature of any component inside a. Dry Bulb Temperature: 55 to 85 °F (13 to the payload fairing with a view factor to the 29 °C) controllable to ±4 °F (±2 °C) of payload with the exception of the Stage 4 motor. setpoint The maximum Stage 4 motor surface temperature b. Dew Point Temperature: 38 to 62 °F (3 to exposed to the payload will not exceed 350 °F 17 °C) (177 °C), assuming no shielding between the aft c. Relative Humidity: determined by drybulb end of the payload and the forward dome of the and dewpoint temperature selections and motor assembly. The Payload Adapter Module generally controlled to within ±15%. (PAM), used with the fairing to provide Relative humidity is bound by the encapsulation of the payload during ground psychrometric chart and will be controlled processing, provides some level of shielding such that the dew point within the fairing is between the payload and Stage 4 motor. Whether never reached. this temperature is attained prior to payload d. Maximum Flow: 500 cfm separation is dependent upon mission timeline.

4.6.2. Powered Flight The fairing peak vent rate is typically less than The maximum fairing inside wall temperature 0.6 psi/sec. Fairing deployment will be initiated at will be maintained at less than 200 °F (93 °C), with a time in flight that the maximum dynamic an emissivity of 0.92 in the region of the payload. pressure is less than 0.01 psf or the maximum free As a non-standard service, a low emissivity molecular heating rate is less than 0.1 BTU/ft2/sec, coating can be applied to reduce emissivity to less as required by the payload.

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4.6.3. Nitrogen Purge (Non-Standard Service) percent. Since the payload processing will be at a If required for spot cooling of a payload GFP facility, it is assumed the Class 10,000 clean component, Orbital will provide GN2 flow to room environment also adhering to these levels of localized regions in the fairing as a non-standard control will be provided by that facility. service. This option is discussed in more detail in Section 8.3.2. Also with the enhanced contamination control option, Orbital provides an Environmental Control 4.7. Payload Contamination Control System (ECS) from payload encapsulation All payload integration procedures, and through vehicle lift-off. The ECS continuously Orbital’s contamination control program have been purges the fairing volume with clean filtered air. designed to minimize the payload’s exposure to Orbital’s ECS incorporates a HEPA filter unit to contamination from the time the payload arrives at provide FED-STD-209 Class M5.5 (10,000) air. the payload processing facility through orbit Orbital monitors the supply air for particulate insertion and separation. The payload is fully matter via a probe installed upstream of the fairing encapsulated within the fairing and Payload inlet duct prior to connecting the air source to the Adapter Module (PAM) at the payload processing payload fairing. facility, assuring the payload environment stays clean in a Class 100,000 environment. All SLV 4.8. Payload Electromagnetic Environment assemblies that affect cleanliness within the The payload Electromagnetic Environment encapsulated payload volume include the fairing (EME) results from two categories of emitters: 1) and the payload cone assembly. These Minotaur IV onboard antennas and, 2) Range assemblies are cleaned such that there is no radar. All power, control and signal lines inside particulate or non-particulate matter visible to the the payload fairing are shielded and properly normal unaided eye when inspected from 2 to 4 terminated to minimize the potential for feet under 50 ft-candle incident light (Visibly Clean Electromagnetic Interference (EMI). The Minotaur Level II). After encapsulation, the fairing envelope IV payload fairing is Radio Frequency (RF) is either sealed or maintained with a positive opaque, which shields the payload from external pressure, Class 100,000 continuous purge of RF signals while the payload is encapsulated. conditioned air. If required, the payload can be provided with enhanced contamination control as Figure 4-12 lists the frequencies and an enhanced option, providing a Class 10,000 maximum radiated signal levels from vehicle environment, low outgassing, and Visibly Clean antennas that are located near the payload during Plus Ultraviolet cleanliness (see Section 8.3.3). ground operations and powered flight. The Provisions exist in the fairing design accommodate specific EME experienced by the payload during dry nitrogen purge as has been demonstrated on ground processing at the PPF and the launch site the Taurus application. will depend somewhat on the specific facilities that are utilized as well as operational details. With the enhanced contamination control However, typically the field strengths experienced option, the Orbital-supplied elements will be by the payload during ground processing with the cleaned and controlled to support a Class 10,000 fairing in place are controlled procedurally and will clean room environment, as defined in Federal be less than 2 V/m from continuous sources and Standard 209. This includes limiting volatile less than 10 V/m from pulse sources. The highest hydrocarbons to maintain hydrocarbon content at EME during powered flight is created by the C- less than 15 ppm and humidity between 35 to 60 Band transponder transmission, which results in

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peak levels at the payload interface plane of This EME should be compared to the payload’s 88 V/m at 5765 MHz (based on Taurus). Range RF susceptibility levels (MIL-STD-461, RS03) to transmitters are typically controlled to provide a define margin. field strength of 10 V/m or less inside the fairing.

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

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5. PAYLOAD INTERFACES payload deflection must be given with the Finite This section describes the available Element Model to evaluate payload dynamic mechanical, electrical and Launch Support deflection with the Coupled Loads Analysis (CLA). Equipment (LSE) interfaces between the Minotaur The payload contractor should assume that the IV launch vehicle and the payload. interface plane is rigid; Orbital has accounted for deflections of the interface plane. The CLA will 5.1. Payload Fairing provide final verification that the payload does not Orbital’s flight-proven Taurus 92-inch diameter violate the dynamic envelope. payload fairing is used to encapsulate the payload, provide protection and contamination control 5.1.2. Payload Access Door during ground handling, integration operations and Orbital provides one 18 in. by 24 in. (45.7 cm flight. The fairing is a bi-conic design made of by 61.0 cm) payload fairing access door to provide graphite/epoxy face sheets with aluminum access to the payload after fairing mate. The door honeycomb core. The two halves of the fairing are can be positioned according to payload structurally joined along their longitudinal interface requirements within the cylindrical section of the using Orbital’s low contamination frangible joint fairing, providing access to the payload without system. An additional circumferential frangible having to remove any portion of the fairing or joint at the base of the fairing supports the fairing break electrical connections. The specific location loads. At separation, a gas pressurization system is defined and controlled in the payload ICD. is activated to pressurize the fairing deployment Additional access doors can readily be provided thrusters. The fairing halves then rotate about as an enhanced option (see Section 8.1.2). external hinges that control the fairing deployment to ensure that payload and launch vehicle 5.2. Payload Mechanical Interface and clearances are maintained. All elements of the Separation System deployment system have been demonstrated Minotaur IV provides for a standard non- through test to comply with stringent separating payload interface and an optional contamination requirements. Orbital-provided payload separation system.

5.1.1. Payload Dynamic Design Envelope The fairing drawing in Figure 5-1 shows the maximum dynamic envelope available for the payload during powered flight. The dynamic envelope shown account for fairing and vehicle structural deflections 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 OSP via the ICD.

No part of the payload may extend aft of the payload interface plane without specific OSP 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 Figure 5-1. Standard 92 in. Fairing Envelope

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Orbital will provide all flight hardware and separation system is utilized, Orbital will provide integration services necessary to attach non- all the wiring through the separable interface separating and separating payloads to Minotaur plane. If the option is not exercised the customer IV. Payload ground handling equipment is will be responsible to provide the wiring from the typically the responsibility of the payload spacecraft to the separation plane. contractor. All attachment hardware, whether Orbital or customer provided, must contain locking 5.3.1. Payload Umbilical Interfaces features consisting of locking nuts, inserts or The payload umbilical connector provides 60 fasteners. wires from the ground to the spacecraft via a dedicated payload umbilical within the vehicle, as 5.2.1. Standard Non-Separating Mechanical shown in Figure 5-3. The length of the internal Interface umbilical is approximately 25 ft (7.62 m). The Orbital’s payload interface design provides a cabling from the LEV to the launch vehicle is standard interface that will accommodate multiple approximately 130 ft (39.6 m). This umbilical is a payload configurations. Figure 5-2 illustrates the dedicated pass through harness for ground standard, non-separating payload mechanical processing support. It allows the payload interface. This is for payloads that provide their command, control, monitor, and power to be easily own separation system or payloads that will not configured per each individual user’s separate. The interface is a standardized circular requirements. The umbilical wiring is configured bolted interface common with the Evolved as a one-to-one from the Payload/Minotaur IV Expendable Launch Vehicle (EELV). The interface through to the payload EGSE interface in interface is a 62.01-inch diameter bolted interface. the Launch Equipment Vault, the closest location A butt joint with 121 holes (0.265-inch diameter) for operating customer supplied payload EGSE designed for ¼-inch fasteners is the payload equipment. mounting surface as shown in Figure 5-2. Alternate or multiple payload configurations can It is a Launch Vehicle requirement that the also be accommodated with the use of a bulkhead payload provide two (2) separation loopback which allows flexibility in mounting patterns and circuits on the payload side of the separation configurations. plane. These are typically wired into different separation connectors for redundancy. These 5.2.2. Orbital Supplied Mechanical Interface breakwires are used for positive separation Control Drawing indication telemetry and initiation of the CCAM Orbital will provide a toleranced Mechanical maneuver. Interface Control Drawing (MICD) to the payload contractor to allow accurate machining of the Figure 5-4 details the pin outs for the standard fastener holes. The Orbital provided MICD is the interface umbilical. All wires are twisted, shielded only approved documentation for drilling the pairs, and pass through the entire umbilical payload interface. system, both vehicle and ground, as one-to-one to simplify and standardize the payload umbilical 5.3. Payload Electrical Interfaces configuration requirements while providing The payload electrical interface supports maximum operational flexibility to the payload battery charging, external power, discrete provider. commands, discrete telemetry, analog telemetry, serial communication, payload separation indications, and up to 16 separate ordnance discretes. If an optional Orbital-provided

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Figure 5-2. Standard, Non-separating Payload Mechanical Interface

Figure 5-3. Payload Electrical Interface Block Diagram, With No Orbital Supplied Separation System

Figure 5-4. Payload 1:1 Umbilical Pin Outs

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5.3.2. Payload Interface Circuitry Standard interface circuitry passing through the payload-to-launch vehicle electrical connections are shown in Figure 5-5. This figure details the interface characteristics for launch vehicle commands, discrete and analog telemetry, separation loopbacks, pyro initiation, and serial communications interfaces 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 battery charger. This power is routed through the payload umbilical cable. Up to 5.0 amperes per wire pair can be accommodated. The payload battery charger should be sized to withstand the line loss from the LEV to the spacecraft.

5.3.4. Payload Command and Control The Minotaur IV standard interface provides discrete sequencing commands generated by the launch vehicle’s Ordnance Driver Module (ODM) that are available to the payload as closed circuit opto-isolator command pulses of 5 A in lengths of 35 ms minimum. The total number of ODM discretes is sixteen (16) and can be used for any combination of (redundant) ordnance events and/or discrete commands depending on the payload requirements. Figure 5-5. Minotaur IV Payload Electrical Interface Block Diagram 5.3.5. Pyrotechnic Initiation Signals 5.3.6. Payload Telemetry Orbital provides the capability to directly The baseline telemetry subsystem capability initiate 16 separate pyrotechnic conductors provides a number of dedicated payload discrete through two dedicated MACH Ordnance Driver (bi-level) and analog telemetry monitors through Modules (ODM). Each ODM provides for up to dedicated channels in the vehicle encoder. Up to eight drivers capable of a 5 A, 100 ms, current 24 channels will be provided with type and data limited pulse into a 1.5 ohm resistive load. All rate being defined in the mission requirements eight channels can be fired simultaneously with an document. In addition, a GCI610 will be utilized in accuracy of 1 ms between channels. In addition, the encoder stack for serial data ranging up to 600 the ODM channels can be utilized to trigger high Kbs if required. The GCI610 utilizes an IRIG impedance discrete events if required. Safing for standard RS422 driver interface for simplicity in all payload ordnance events will be accomplished payload interface definition. The payload serial through an Arm/Disarm (A/D) Switch. and analog data will be embedded in the baseline

vehicle telemetry format. For discrete monitors,

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the payload customer must provide the 5 Vdc (programmable) and data transmission bit rates. source and the return path. The current at the The number of channels, sample rates, etc. will be payload interface must be less than 10 mA. defined in the Payload ICD. Separation breakwire monitors can be specified if required. The number of analog channels 5.3.9. Non-Standard Electrical Interfaces available for payload telemetry monitoring is Non-standard services such as serial dependent on the frequency of the data. Payload command and telemetry interfaces can be telemetry requirements and signal characteristics negotiated between OSP and the payload will be specified in the Payload ICD and should not contractor on a mission-by-mission basis. The change once the final telemetry format is released selection of the separation system could also at approximately L-6 months. Orbital will tape, impact the payload interface design and will be archive, and reduce the data into a digital format defined in the Payload ICD. for delivery to the payloaders for review. 5.3.10. Electrical Launch Support Equipment Due to the use of strategic assets, Minotaur IV Orbital will provide space for a rack of telemetry is subject to the provisions of the customer supplied EGSE in the LCR, or either of Strategic Arms Reduction Treaty (START). the on-pad equipment vaults. The equipment will START treaty provisions require that certain interface with the launch vehicle/spacecraft Minotaur IV telemetry be unencrypted and through either the dedicated payload umbilical provided to the START treaty office for interface or directly through the payload access dissemination to the signatories of the treaty. The door. The payload customer is responsible for extent to which START applies to the payload providing cabling from the EGSE location to the telemetry will be determined by RSLP. launch vehicle/spacecraft. Encrypted payload telemetry can be added as a Separate payload ground processing non-standard service pending approval by RSLP and the START treaty office. harnesses that mate directly with the payload can be accommodated through the payload access 5.3.7. Payload Separation Monitor Loopbacks door(s) as defined in the Payload ICD. Separation breakwire monitors are required on 5.4. Payload Design Constraints both sides of the payload separation plane. With the Orbital provided separation systems, Minotaur The following sections provide design IV provides three (3) separation loopbacks on the constraints to ensure payload compatibility with the Minotaur IV system. launch vehicle side of the separation plane for positive payload separation indication. 5.4.1. Payload Center of Mass Constraints

Minotaur IV also requires two (2) separate Along the Y and Z-axes, the payload CG must loopbacks on the payload side of the separation be within 1.0 inch (3.8 cm) of the vehicle centerline. Payloads whose CG extend beyond plane. These are used for telemetry indication of separation and also the initiation of the Stage 4 the 1.0 inch lateral offset limit will require Orbital to CCAM maneuver. verify the specific offsets that can be accommodated. 5.3.8. Telemetry Interfaces 5.4.2. Final Mass Properties Accuracy The standard Minotaur IV payload interface provides a 16Kbps RS-422/RS-485 serial interface The final mass properties statement must for payload use with the flexibility to support a specify payload weight to an accuracy of at least 1 lbm (0.5 kg), the center of gravity to an accuracy variety of channel/bit rate requirements, and provide signal conditioning, PCM formatting of at least 0.25 inch (6.4 mm) in each axis, and the

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products of inertia to an accuracy of at least 0.5 must schedule all RF tests at the integration site slug-ft2 (0.7 kg-m2) (see Figure 5-6). In addition, with Orbital in order to obtain proper range if the payload uses liquid propellant, the slosh clearances and protection. frequency must be provided to an accuracy of 0.2 Hz, along with a summary of the method used to 5.4.5. Payload Dynamic Frequencies determine slosh frequency. To avoid dynamic coupling of the payload modes with the natural frequency of the vehicle, the spacecraft should be designed with a structural stiffness to ensure that the lateral fundamental frequency of the spacecraft, fixed at the spacecraft interface is typically greater than 25 Hz (based on Taurus). However, this value is effected significantly by other factors such as incorporation of a spacecraft isolation system and/or separation system. Therefore, the final

Figure 5-6. Payload Mass Properties determination of compatibility must be made on a Measurement Tolerance mission-specific basis.

5.4.3. Pre-Launch Electrical Constraints 5.4.6. Payload Propellant Slosh Prior to launch, all payload electrical interface A slosh model should be provided to Orbital in circuits are constrained to ensure there is no either the pendulum or spring-mass format. Data current flow greater than 10 mA across the on first sloshing mode are required and data on payload electrical interface plane. The primary higher order modes are desirable. The slosh support structure of the spacecraft shall be model should be provided with the payload finite electrically conductive to establish a single point element model submittals. electrical ground. 5.4.7. Payload-Supplied Separation Systems 5.4.4. Payload EMI/EMC Constraints If the payload employs a non-Orbital separation system, then the shock delivered to the The Minotaur IV avionics share the payload Stage 4 vehicle interface must not exceed the limit area inside the fairing such that radiated level characterized in Section 4.3 (Figure 4-6). As emissions compatibility is paramount. OSP places stated in that section, shock above this level could no firm radiated emissions limits on the payload require a requalification of components or an other than the prohibition against RF acceptance of risk by RSLP. transmissions within the payload fairing. Prior to launch, Orbital requires review of the payload 5.4.8. System Safety Constraints radiated emission levels (MIL-STD-461, RE02) to OSP considers the safety of personnel and verify overall launch vehicle EMI safety margin equipment to be of paramount importance. EWR (emission) in accordance with MIL-E-6051. 127-1 outlines the safety design criteria for Payload RF transmissions are not permitted after Minotaur IV payloads. These are compliance fairing mate and prior to an ICD specified time documents and must be strictly followed. It is the after separation of the payload. An EMI/EMC responsibility of the customer to ensure that the analysis may be required to ensure RF payload meets all OSP, Orbital, and range compatibility. imposed safety standards.

Payload RF transmission frequencies must be Customers designing payloads that employ coordinated with Orbital and range officials to hazardous subsystems are advised to contact ensure non-interference with Minotaur IV and OSP early in the design process to verify compliance with system safety standards. range transmissions. Additionally, the customer

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6. MISSION INTEGRATION coordinate all mission planning and contracting activities. RSLP is supported by Northrop 6.1. Mission Management Approach Grumman and other associate contractors for The Minotaur IV program is managed through technical and logistical support, particularly U.S. Air Force, Space and Missile Systems utilizing their extensive expertise and background Center, Rocket Systems Launch Program (RSLP). knowledge of the Peacekeeper booster and RSLP serves as the primary point of contact for subsystems. the payload customers for the Minotaur IV launch service. A typical integrated OSP organizational 6.1.2. Orbital Mission Responsibilities structure is shown in Figure 6-1. Open As the launch vehicle provider, Orbital’s communication between RSLP , Orbital, and the responsibilities fall into four primary areas: customer, emphasizing timely transfer of data and a. Launch Vehicle Program Management prudent decision-making, ensures efficient launch b. Mission Management vehicle/payload integration operations. c. Engineering d. Launch Site Operations 6.1.1. RSLP Mission Responsibilities The program office for all OSP missions is the Orbital assigns a Mission Manager to manage RSLP . They are the primary Point of Contact the launch vehicle technical and programmatic (POC) for all contractual and technical interfaces for a particular mission. The Orbital coordination. RSLP contracts with Orbital to Mission Manager is the single POC for all aspects provide the Launch Vehicle and launch integration of a specific mission. This person has overall and separately with commercial Spaceports and/or program authority and responsibility to ensure that Government Launch Ranges for launch site payload requirements are met and that the facilities and services. Once a mission is appropriate launch vehicle services are provided. identified, RSLP will assign a Mission Manager to The Orbital Mission Manager will jointly chair the

Figure 6-1. Typical Integrated OSP Organizational Structure

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Mission Integration Working Groups (MIWGs) with d. Range interface, safety, and flight the RSLP Mission Manager. The Mission operations activities, document Managers responsibilities include detailed mission exchanges, meetings and reviews. planning, payload integration services, systems engineering, mission-peculiar design and analyses Figure 6-2 details the typical Mission Cycle for coordination, payload interface definition, launch a specific launch and how this cycle folds into the range coordination, integrated scheduling, launch Orbital vehicle production schedule with typical site processing, and flight operations. payload activities and milestones. A typical Mission Cycle is based on an 18 month interval 6.2. Mission Planning and Development between mission authorization and launch. This OSP will assist the customer with mission interval reflects the OSP contractual schedule and planning and development associated with has been shown to be an efficient schedule based Minotaur IV launch vehicle systems. These on Orbital’s Minotaur, Taurus and Pegasus services include interface design and configuration program experience. However, OSP is flexible to control, development of integration processes, negotiate either accelerated cycles, which take launch vehicle analyses, facilities planning, launch advantage of the Minotaur IV/Pegasus/Minotaur/ campaign planning to include range services and Taurus multi-customer production sets, or special operations, and integrated schedules. extended cycles required by unusual payload requirements, such as extensive analysis or The procurement, analysis, integration and complex payload-launch vehicle integrated test activities required to place a customer’s designs or tests or funding limitations. payload into orbit are typically conducted over a 20 month standard sequence of events called the 6.3. Mission Integration Process Mission Cycle. This cycle normally begins 18 months before launch, and extends to 8 weeks 6.3.1. Integration Meetings after launch. The core of the mission integration process consists of a series of Mission Integration and Once contract authority to proceed is received, Range Working Groups (MIWG and RWG, the Mission Cycle is initiated. The contract option respectively). The MIWG has responsibility for all designates the payload, launch date, and basic physical interfaces between the payload and the mission parameters. In response, the Minotaur IV launch vehicle. As such, the MIWG creates and Program Manager designates an Orbital Mission implements the Payload-to-Minotaur IV ICD in Manager who ensures that the launch service is addition to all mission-unique analyses, hardware, supplied efficiently, reliably, and on-schedule. software, and integrated procedures. The RWG is responsible for the areas of launch site operations; The typical Mission Cycle interweaves the range interfaces; safety review and approval; and following activities: flight design, trajectory, and guidance. a. Mission management, document Documentation produced by the RWG includes all exchanges, meetings, and formal reviews required range and safety submittals. required to coordinate and manage the launch service. Working Group membership consists of the b. Mission analyses and payload integration, Mission Manager and representatives from document exchanges, and meetings. Minotaur IV engineering and operations c. Design, review, procurement, testing and organizations, as well as their counterparts from integration of all mission-peculiar the customer organization. While the number of hardware and software. meetings, both formal and informal, required to develop and implement the mission integration

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Figure 6-2. Typical Mission Integration Schedule process will vary with the complexity of the variability in complexity of different payloads and spacecraft, quarterly meetings are typical. missions, the content and number of these reviews can be tailored to customer requirements. 6.3.2. Mission Design Reviews (MDR) As a baseline, Orbital will conduct two readiness Two mission-specific design reviews will be reviews as described below. held to determine the status and adequacy of the a. Mission Readiness Review — launch vehicle mission preparations. They are Conducted within 1 month of launch, the designated MDR-1 and MDR-2 and are typically Mission Readiness Review (MRR) held 6 months and 13 months, respectively, after provides a pre-launch assessment of authority to proceed. They are each analogous to integrated launch vehicle/payload/facility Preliminary Design Reviews (PDRs) and Critical readiness prior to committing significant Design Reviews (CDRs), but focus primarily on resources to the launch campaign. mission-specific elements of the launch vehicle b. Launch Readiness Review — The effort. Launch Readiness Review (LRR) is conducted at L-1 day and serves as the 6.3.3. Readiness Reviews final assessment of mission readiness During the integration process, reviews are prior to activation of range resources on held to provide the coordination of mission the day of launch. participants and management outside of the regular contact of the Working Groups. Due to the

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6.4. Documentation analyses. Preliminary mass properties should be Integration of the payload requires detailed, submitted as part of the MRD at launch vehicle complete, and timely preparation and submittal of authority to proceed. Updated mass properties interface documentation. As the launch service shall be provided at predefined intervals identified provider, RSLP is the primary communication path during the initial mission integration process. with support agencies, which include—but are not Typical timing of these deliveries is included in limited to—the various Range support agencies Figure 6-2. and U.S. Government agencies such as the U.S. Department of Transportation and U.S. State 6.4.1.3. Payload Finite Element Model Department. Customer-provided documents A payload mathematical model is required for represent the formal communication of use in Orbital’s preliminary coupled loads requirements, safety data, system descriptions, analyses. Acceptable forms include either a and mission operations planning. The major Craig-Bampton model valid to 120 Hz or a products and submittal times associated with NASTRAN finite element model. For the final these organizations are divided into two areas— coupled loads analysis, a test verified those products that are provided by the customer, mathematical model is desired. and those produced by Orbital. 6.4.1.4. Payload Thermal Model for Integrated 6.4.1. Customer-Provided Documentation Thermal Analysis Documentation produced by the customer is An integrated thermal analysis can be detailed in the following paragraphs. performed for any payload as a non-standard service. A payload thermal model will be required 6.4.1.1. Payload Questionnaire from the payload organization for use in Orbital’s The Payload Questionnaire is designed to integrated thermal analysis if it is required. The provide the initial definition of payload analysis is conducted for three mission phases: requirements, interface details, launch site a. Prelaunch ground operations; facilities, and preliminary safety data to OSP. The b. Ascent from lift-off until fairing jettison; and customer shall provide a response to the Payload c. Fairing jettison through payload Questionnaire form (Appendix A), or provide the deployment. same information in a different format, in time to support the Mission Kickoff Meeting. The Models must be provided in SINDA format. customer’s responses to the payload There is no limit on model size although turn- questionnaire define the most current payload around time may be increased for large models. requirements and interfaces and are instrumental in Orbital’s preparation of numerous documents 6.4.1.5. Payload Drawings including the ICD, Preliminary Mission Analysis, Orbital prefers electronic versions of payload and launch range documentation. Additional configuration drawings to be used in the mission pertinent information, as well as preliminary specific interface control drawing, if possible. payload drawings, should also be included with Orbital will work with the customer to define the the response. Orbital understands that a definitive content and desired format for the drawings. response to some questions may not be feasible. These items are defined during the normal mission 6.4.1.6. Program Requirements Document integration process. (PRD) Mission Specific Annex Inputs To obtain range support, a PRD must be 6.4.1.2. Payload Mass Properties prepared. This document describes requirements Payload mass properties must be provided in needed to generally support the Minotaur IV a timely manner in order to support efficient launch launch vehicle. For each launch, an annex is vehicle trajectory development and dynamic submitted to specify the range support needed to

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meet the mission’s requirements. This annex Before a spacecraft arrives at the processing includes all payload requirements as well as any site, the payload organization must provide the additional Minotaur IV requirements that may arise cognizant range safety office with certification that to support a particular mission. The customer the system has been designed and tested in completes all appropriate PRD forms for submittal accordance with applicable safety requirements to Orbital. (e.g. EWR 127-1 Range Safety Requirements for baseline and VAFB Payload Integration missions). 6.4.1.6.1. Launch Operations Requirements Spacecraft that integrate and/or launch at a site (OR) Inputs different than the processing site must also comply To obtain range support for the launch with the specific launch site’s safety requirements. operation and associated rehearsals, an OR must Orbital will provide the customer coordination and be prepared. The customer must provide all guidance regarding applicable safety payload pre-launch and launch day requirements requirements. for incorporation into the mission OR. It cannot be overstressed that the applicable 6.5. Safety safety requirements should be considered in the earliest stages of spacecraft design. Processing 6.5.1. System Safety Requirements and launch site ranges discourage the use of In the initial phases of the mission integration waivers and variances. Furthermore, approval of effort, regulations and instructions that apply to such waivers cannot be guaranteed. spacecraft design and processing are reviewed. Not all safety regulations will apply to a particular 6.5.2. System Safety Documentation mission integration activity. Tailoring the range For each Minotaur IV mission, OSP acts as requirements to the mission unique activities will the interface between the mission and Range be the first step in establishing the safety plan. Safety. In order to fulfill this role, OSP requires OSP has three distinctly different mission safety information from the payloader. For approaches affecting the establishment of the launches from either the Eastern or Western safety requirements: Ranges, EWR 127-1 provides detailed range a. Baseline mission: Payload integration and safety regulations. To obtain approval to use the launch operations are conducted at VAFB, launch site facilities, specified data must be CA prepared and submitted to the OSP Program b. Campaign/VAFB Payload Integration Office. This information includes a description of mission: Payload integration is conducted each payload hazardous system and evidence of at VAFB and launch operations are compliance with safety requirements for each conducted from a non-VAFB launch system. Drawings, schematics, and assembly and location. handling procedures, including proof test data for c. Campaign/Non-VAFB Payload Integration all lifting equipment, as well as any other mission: Payload integration and launch information that will aid in assessing the respective systems should be included. Major categories of operations are conducted at a site other hazardous systems are ordnance devices, than VAFB. radioactive materials, propellants, pressurized

systems, toxic materials, cryogenics, and RF For the baseline and VAFB Payload radiation. Procedures relating to these systems as Integration missions, spacecraft prelaunch well as any procedures relating to lifting operations operations are conducted at Government or battery operations should be prepared for safety Furnished Property (GFP) Payload Processing review submittal. OSP will provide this information Facility (PPF). For campaign style missions, the to the appropriate safety offices for approval. spacecraft prelaunch operations are performed at the desired launch site.

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Release 1.1 January 2006 48 Minotaur IV User’s Guide Section 7.0 – Ground and Launch Operations

7. GROUND AND LAUNCH OPERATIONS the master document communicating all activities planned at the field site. The schedule contains 7.1. Minotaur IV/Payload Integration Overview notations regarding the status of the work package The processing of the Minotaur IV utilizes document and hardware required to begin the many of the same proven techniques developed operation. Mission-specific work packages are for the Pegasus, Taurus and Minotaur launch created for mission-unique or payload-specific vehicles. This minimizes the handling complexity procedures. Any discrepancies encountered are for both vehicle and payload. recorded on a Discrepancy Report and dispositioned as required. All activities are in 7.2. Ground And Launch Operations accordance with Orbital’s ISO 9001 certification. Ground and launch operations are conducted in three major phases: 7.2.1.2. GCA/Orion 38 Integration and Test a. Launch Vehicle Integration — Assembly Activities and test of the Minotaur IV vehicle The GCA will undergo subsystem level testing b. Payload Processing/Integration — at Orbital’s Chandler facility prior to being shipped Receipt and checkout of the satellite to the field site. The GCA and the Stage 4 Orion payload, followed by integration with 38 motor are then delivered to the launch vehicle Minotaur IV fairing and Payload Adapter processing facility located at VAFB. Upon arrival Module (PAM) and verification of at VAFB these components/sub-assemblies will interfaces undergo a thorough inspection and subsystem c. Launch Operations — Includes transport level checkout. At this time range certification of to the launch pad, final integration, Range Tracking System (RTS) and Flight checkout, arming and launch. Termination System (FTS) devices will be performed. The components will be reinstalled 7.2.1. Launch Vehicle Integration and in-vehicle testing of the RTS and FTS Orbital will process all Minotaur IV vehicles systems will be performed. After the completion of according to a flow similar to that implemented for subsystem level testing the Orion 38 motor is the Minotaur and Taurus vehicles. All launch integrated into the GCA to form the Stage 4 vehicle motors, parts and completed assembly. subassemblies are delivered to the launch Vehicle Processing Facility (VPF) from Orbital’s Chandler 7.2.1.3. PK Motor Integration and Test production facility, the assembly/motor vendor or Activities the Government. Figure 7-1 depicts the typical The PK Stage 1, 2 and 3 motors are delivered flow of hardware from the factory to the launch to the launch vehicle processing facility where they site. Flowcharts of the field processing are shown undergo checkout and testing. Once integration is in Figure 7-2. complete, a booster confidence test will be conducted. 7.2.1.1. Planning and Documentation

Minotaur IV integration and test activities are 7.2.1.4. Mission Simulation Tests controlled by a comprehensive set of Work Orbital will run three Mission Simulator Tests Packages (WPs) that describe and document (MST) to verify the functionality of launch vehicle every aspect of integrating and testing Minotaur IV hardware, and software. The Mission Simulation and its payload. All testing and integration Tests use the actual flight software and simulate a activities are scheduled by work package number “fly to orbit” scenario using simulated Inertial on a daily activity schedule updated and Navigation System (INS) data. This will allow the distributed daily during field operations. This test to proceed throughout all mission phases schedule is maintained by Orbital and serves as

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

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Figure 7-2. SLV Processing Flow recording vehicle performance data. The data will installation of the launch stool. This stool is the be compared to simulations performed in the same design used for Orbital’s Taurus SLV. It factory software laboratory using an identical copy supports a flat pad launch of a full PK booster of the flight software. Orbital will use GFP PK assembly with front section and fourth stage motor nozzle assembly simulators to perform all mission options. After stool installation, the fixed simulations. These components will provide a scaffolding installation is performed. This realistic assessment of booster performance scaffolding provides access to the base ring of the during the testing operations. After a thorough PK Stage 1 motor during integration activities. data review of all telemetry parameters, the test configuration is disassembled and setup for Once the booster arrives at the launch site, it payload integration begins. is then lifted and emplaced onto the launch stool. Each motor assembly will be individually stacked 7.2.1.5. Booster Assembly Stacking/Launch using a process developed for handling Taurus Pad Preparation Stage 0 motors. Scaffolding integration continues After completion of the MST, the booster as the booster stages are mated. assembly (Stages 1, 2 and 3) and the stage 4 assembly (Orion 38 integrated with the GCA) are The Stage 4 assembly is shipped in the transported to the launch facility. Figure 7-3 vertical configuration to the launch facility for shows a pictorial representation of the processing payload integration. flow. 7.2.2. Payload Processing/Integration Prior to the arrival of the PK boosters, the site Payloads normally undergo initial checkout is prepared for launch operations with the and preparation for launch at an Air Force payload

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Figure 7-3. Minotaur IV Processing Flow processing facility (PPF) or commercial facilities at loading facilities in the VAB. This is a non- VAFB. After arrival at the PPF, the payload standard service. completes its own independent verification and checkout prior to beginning integrated processing 7.2.2.2. Final Vehicle Integration and Test with Minotaur IV fairing and Payload Adapter After successful completion of payload Module (PAM). The Minotaur IV fairing and PAM mate/fairing closeout the completed front section will be delivered to the payload processing facility assembly (Minotaur IV Stage 4 assembly for encapsulation of the payload. The fairing and integrated with the payload assembly) will then be PAM provide a sealed enclosure which protects lifted in vertical configuration atop the booster the payload and provides a structure to facilitate assembly. Figure 7-4 illustrates the vertical lifting transportation to the launch facility. After operation performed on a Taurus front section. enclosure of the payload in the fairing, the Final post mates checks of the booster assembly assembly is shipped in the vertical configuration to and front section assembly interface are then the launch facility for a pre-installation verification conducted. A final systems verification test, test. similar to the previous MST, is then performed. At this point the vehicle is ready for final Range 7.2.2.1. Payload Propellant Loading interface tests and launch readiness. Payloads utilizing integral propulsion systems with propellants such as hydrazine can be loaded and secured through coordinated Orbital and contractor arrangements for use of the propellant

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7.3. Launch Operations

7.3.1. Launch Control Organization The Launch Control Organization is split into two groups: the Management group and the Technical group. The Management group consists of senior range personnel and Mission Directors/Managers for the launch vehicle and payload. The Technical Group consists of the personnel responsible for the execution of the launch operation and data review/assessment for the Payload, the Launch Vehicle and the Range. The Payload’s members of the technical group are engineers who provide technical representation in the control center. The Launch Vehicle’s members of the technical group are engineers who prepare the Minotaur IV 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 Figure 7-4. Minotaur IV Upper Stack Assembly maintain and monitor the voice and data will be Vertically Integrated to Minotaur IV equipment, tracking facilities and all assets Booster Assembly in a Similar Manner to involved with RF communications with the launch Taurus Upper Stack vehicle. In addition, the Range provides personnel responsible for the Flight Termination System monitoring and commanding.

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Release 1.1 January 2006 54 Minotaur IV User’s Guide Section 8.0 – Optional Enhanced Capabilities

8. OPTIONAL ENHANCED CAPABILITIES the separation system to the payload is allocated The OSP launch service is structured to to the separation system and included in the provide a baseline vehicle configuration which is launch vehicle mass. then augmented with optional enhancements to meet the unique needs of individual payloads. Separation velocity is provided by up to eight The baseline vehicle capabilities are defined in the matched spring actuators. The spring assemblies previous sections and the optional enhanced may be tailored to mitigate the effects of payload capabilities are defined below. The enhanced CG offset, controlling tip-off within 5 deg/sec. Tip- options allow customization of launch support and off rates are highly dependent on payload mass accommodations to the Minotaur IV designs on an properties, but are typically on the order 1 efficient, “as needed” basis. deg/sec. Preliminary and final mission-specific tip- off analyses are conducted for each payload using 8.1. Mechanical Interface and Separation Orbital’s computer simulation dynamic analysis System Enhancements tools. If non-standard separation velocities are needed, different springs may be substituted on a 8.1.1. Separation Systems mission-specific basis as a non-standard service. Various separation systems can be provided or accommodated to meet mission-unique Other separation systems can also be requirements. As a baseline option, Orbital offers supplied on a mission-specific basis. an optional payload separation system that is flight proven on Taurus. The separation system is 8.1.2. Additional Fairing Access Doors manufactured for Orbital by SAAB Ericson Space Additional access doors can be provided to (SES). SES has extensive experience in supplying accommodate unique payload requirements. The separation systems for a wide range of launch standard door size is 18 in. by 24 in. (45.7 cm by vehicles and payloads. This system is based on a 61.0 cm). Access doors of non-standard size can design that has flown over 30 times with 100% also be provided as necessary. Orbital performs success. The baseline separation system, shown structural analyses to verify the acceptability of the in Figure 8-1, has a standardized 38.81 inch bolt mission-specific door configuration. Other fairing pattern. It is a marmon clamp design employing access configurations, such as small circular two aluminum interface rings that are clamped by access panels, can also be provided as negotiated dual, semi-circular stainless steel clamp bands mission-specific enhancements. with aluminum clamp shoes. Each of the two retention bolts is severed by a redundantly 8.1.3. Payload Isolation System initiated bolt cutter. OSP offers a flight-proven payload isolation system as a non-standard service. The Softride The separation ring to which the payload for Small Satellites (SRSS) was developed by Air attaches is supplied with through holes and the Force Research Laboratory (AFRL) and CSA separation system is mated to the spacecraft Engineering. It was successfully demonstrated on during processing at the PPF. The weight of the two initial Minotaur missions and six Taurus hardware separated with the payload is missions. The typical Minotaur configuration is approximately 8.7 lbm (13.95 kg). Orbital- shown in Figure 8-2. This mechanical isolation provided attachment bolts to this interface can be system has demonstrated the capability to inserted from either the launch vehicle or the significantly alleviate the transient dynamic loads payload side of the interface (NAS630xU, dash that occur during flight. The isolation system can number based on payload flange thickness). The provide relief to both the overall payload center of weight of the bolts, nuts, and washers connecting gravity loads and component or subsystem

Release 1.1 January 2006 55 Minotaur IV User’s Guide Section 8.0 – Optional Enhanced Capabilities

Figure 8-1. 38-in. Separation System Option responses. Typically the system will reduce 8.2. Performance Enhancements transient loads to approximately 50% of the level they would be without the system. The exact 8.2.1. Insertion Accuracy results can be expected to vary for each particular Enhanced insertion accuracy or support for spacecraft and with location on the spacecraft. multiple payload insertion can be provided as an Generally, a beneficial reduction in shock and enhanced option utilizing the Hydrazine Auxiliary vibration will also be provided. The isolation Propulsion System (HAPS). The common usage system does impact overall vehicle performance of the Orion 38 makes the flight proven HAPS (by approximately 20 to 40 lb [9 to 18 kg]) and the design directly applicable to the Minotaur IV available payload dynamic envelope by up to 4 usage, as shown in Figure 8-3. Orbital insertion inches (10.16 cm) axially and up to 1.0 inch (2.54 accuracy can typically be improved to ±10 nmi cm) laterally. (±18.5 km) or better in each apse and ±0.05 deg in inclination. For orbits above 324 nmi (600 km), the

Release 1.1 January 2006 56 Minotaur IV User’s Guide Section 8.0 – Optional Enhanced Capabilities

HAPS also permits injection of shared payloads into different orbits. HAPS, which is mounted inside the Avionics Structure, consists of a hydrazine propulsion subsystem and a Stage 4 separation subsystem. After burnout and separation from the Stage 4 motor, the HAPS hydrazine thrusters provide additional velocity and both improved performance and precise orbit injection. The HAPS propulsion subsystem consists of a centrally mounted tank containing approximately 130 lbm (59 kg) of hydrazine, helium pressurization gas, and three fixed, axially pointed thrusters. The hydrazine tank contains an Figure 8-2. Soft Ride Payload Isolation System integral bladder which will support multiple as Integrated on Minotaur LV restarts.

HAPS can also increase payload mass by 8.2.2. Star 48 Stage 4 approximately 50 to 250 lbm (22.7 to 113 kg), The modular design of Orbital’s GCA and depending on the orbit. Specific performance integrating structures provides great flexibility in capability associated with the HAPS can be accommodating alternative Stage 4 propulsion provided by contacting the OSP program office. systems. As one low risk example, a optional

Figure 8-3. Hydrazine Auxiliary Propulsion System (HAPS) Used to Provide Insertion Accuracy

Release 1.1 January 2006 57 Minotaur IV User’s Guide Section 8.0 – Optional Enhanced Capabilities

configuration using the more powerful ATK Thiokol 100,000 or Class 10,000). Nitrogen purge is used Star-48 motor is shown in Figure 8-4. The only in conjunction with the Conditioned Air option to modifications required to accommodate this provide mission-specific localized cooling and/or change are a modified Motor Adapter Cone (MAC) dry nitrogen environments to satisfy unique with the Star 48 forward interface and a longer 3/4 payload environmental requirements. interstage to allow room for the increased motor length. Other alternative motors can be similarly 8.3.2. Nitrogen Purge adopted. Continuous clean dry nitrogen inside the shroud during vehicle processing from payload encapsulation to launch is available as an option. Dry clean nitrogen purge can be provided to the payload at a Class 10,000 environment for continuous purge of the payload after fairing encapsulation until lift-off. The capability was demonstrated on the Minotaur MightySat mission with the exception of purge during transportation.

A nitrogen cooling system is already provided on every SLV mission to spot-cool sensitive electronic boxes. Flow adjustments for cooling versus purge would be changed back and forth to accommodate both. If nitrogen purge is required during transport, the only additional items needed would be a minor addition of a nitrogen bottle, regulator, and mounting hardware.

The system distribution lines are routed across Figure 8-4. Orion 38 Stage 4 Motor can be the payload interface plane and/or along the inner Replaced with a Star-48 to Provide Increased surface of the shroud or fairing. If required for Performance spot cooling of a payload component, Orbital will

provide GN2 flow to localized regions in the 8.3. Environmental Control Options fairing. The GN2 will meet Grade B specifications, as defined in MIL-P-27401C and can be regulated 8.3.1. Conditioned Air to at least 5 scfm. The system’s regulators are set Conditioned air can be provided within the to a desired flow rate during pre-launch fairing volume using an Environmental Control processing. Payload purge requirements are System (ECS) via a “fly-out” duct that is retracted controlled and documented via the launch vehicle at launch. Temperature and humidity is regulated to payload ICD. within the limits specified in the Payload ICD. A filter is installed to provide a Class 100,000 Payload purge requirements must be environment, typically. The Nitrogen Purge coordinated with Orbital via the ICD to ensure that (Section 4.6.3) and Enhanced Contamination the requirements can be achieved. Control (Section 4.7 and 8.3.3) enhancements complement this capability. Upon exercise of the 8.3.3. Enhanced Contamination Control Enhanced Cleanliness option, a certified HEPA Understanding that some payloads have filter is used in the input duct to assure the requirements for enhanced cleanliness, OSP necessary low particulate environment (Class offers a contamination control option, which is

Release 1.1 January 2006 58 Minotaur IV User’s Guide Section 8.0 – Optional Enhanced Capabilities

composed of the elements in the following upstream of the fairing inlet duct prior to sections (which is also discussed in Section 4.7). connecting the air source to the payload fairing. Minotaur IV customers can also coordinate combinations of the elements listed below to meet 8.3.4. Launch Pad Environmental Control the unique needs of their payloads. For launch sites without gantries for environmental control or vehicle access, optional 8.3.3.1. High Cleanliness Integration Quick Erect Scaffold® (QES) will protect the Environment (Class 10K or 100K) vehicle from the environments, maintain With enhanced contamination control, a soft temperatures within 60 to 100 °F (in conjunction walled clean room can be provided to ensure a with a thermal blanket) and provide access to the FED-STD-209 Class M6.5 (100,000) or Class vehicle for launch pad operations (see Figure 8-5). M5.5 (10,000) environment during all payload As the name implies, QES can be rapidly processing activities up to fairing encapsulation. assembled and is highly adaptable for The soft walled clean room and anteroom(s) utilize accommodating different vehicle configurations. HEPA filter units to filter the air and hydrocarbon This scaffolding has been previously content is maintained at 15 ppm or less. The demonstrated on Minotaur. payload organization is responsible for providing the necessary clean room garments for payload 8.3.4.1. Booster Temperature Control staff as well as vehicle staff that need to work The thermal blanket design successfully used inside the clean room. during Minotaur missions can be used to maintain the PK booster operating temperature within the 8.3.3.2. Fairing Surface Cleanliness Options limits of 60 to 100 °F. The thermal blanket is The inner surface of the fairing and payload cone assemblies can be cleaned to cleanliness criteria which ensures no particulate matter visible with normal vision when inspected from 6 to 18 inches under 100 ft-candle incident light. The same will be true when the surface is illuminated using black light, 3200 to 3800 Angstroms (Visibly Clean Plus Ultraviolet). In addition, Orbital can ensure that all materials used within the encapsulated volume have outgassing characteristics of less than 1.0% TML and less than 0.1% CVCM. Items that do not meet these levels can be masked to ensure they are encapsulated and will have no significant effect on the payload.

8.3.3.3. High Cleanliness Fairing Environment With the enhanced contamination control option, Orbital provides an ECS from payload encapsulation until just prior to vehicle lift-off. The ECS continuously purges the fairing volume with clean filtered air. Orbital’s ECS incorporates a HEPA filter unit to provide FED-STD-209 Class Figure 8-5. Mobile Scaffolding for M5.5 (10,000) air. Orbital monitors the supply air Environmental Control Demonstrated on for particulate matter via a probe installed Minotaur Missions

Release 1.1 January 2006 59 Minotaur IV User’s Guide Section 8.0 – Optional Enhanced Capabilities

constructed of outer PVC material with an inner provides better than 100 m position accuracy with insulating liner. It is a four piece cover with Velcro 10 Hz data rate. This capability was successfully seams running along the length of the boosters. demonstrated on the inaugural Minotaur mission. Integral inflatable manifold tubes space the blanket away from the booster and provide space 8.5. Shared Launch Accommodations for ducting conditioned air for the boosters. The Minotaur IV is uniquely capable of providing baseline blanket design only covers the Stage 1, 2 launches of multiple satellite payloads, leveraging and 3 boosters. RSLP and Orbital’s extensive experience in integrating and launching multiple payloads. 8.4. Enhanced Telemetry Options Multiple spacecraft configurations have been flown OSP can provide mission specific on many of Orbital’s Pegasus, Taurus and instrumentation and telemetry components to Minotaur missions to date. A number of different structural configurations have been developed for support additional payload or experiment data dual payloads, one is shown in Figure 8-6. acquisition requirements. Telemetry options Because of the modular nature of the structures, include additional payload-dedicated bandwidth dual payload configurations can be easily and GPS-based precision navigation data. accommodated by the Minotaur IV structural

design. 8.4.1. Enhanced Telemetry Bandwidth

Enhanced mission specific instrumentation and telemetry can be provided, supplying a dedicated telemetry link to support additional payload or experiment data acquisition requirements. A baseline data rate of 1 Mbps is available, however, maximum data rates depend on the mission coverage required and the launch range receiver characteristics and configuration. The enhanced telemetry option was demonstrated on both inaugural Minotaur TLV and SLV missions.

8.4.2. Enhanced Telemetry Instrumentation To support the higher data rate capability in Section 8.4.1, enhanced telemetry instrumentation can be provided. The instrumentation can include strain gauges, temperature sensors, accelerometers, analog data, and digital data configured to mission-specific requirements. This capability was successfully demonstrated on the inaugural OSP-SLV mission. Figure 8-6. Modular Minotaur IV Structural Design Easily Accommodates Multiple 8.4.3. Navigation Data Payloads Precision navigation data using an independent Global Positioning System (GPS) receiver and telemetry link is available as an enhanced option. This option utilizes Orbital’s flight proven GPS Position Beacon (GPB) to provide missile state data for range safety and

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APPENDIX A

PAYLOAD QUESTIONNAIRE

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SATELLITE IDENTIFICATION

FULL NAME:

ACRONYM:

OWNER/OPERATOR:

INTEGRATOR(s):

ORBIT INSERTION REQUIREMENTS*

SPHEROID Standard (WGS-84, Re = 6378.137 km) Other: ALTITUDE Insertion Apse: Opposite Apse: nmi nmi ___ ± __ km ± km or... Semi-Major Axis: Eccentricity: nmi ___ ± __ km ≤ e ≤ INCLINATION ± deg ORIENTATION Argument of Perigee: Longitude of Ascending Node (LAN):

± deg ± deg Right Ascension of Ascending Node (RAAN):

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

LAUNCH WINDOW REQUIREMENTS

NOMINAL LAUNCH DATE:

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

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GROUND SUPPORT EQUIPMENT

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

SEB Describe (in the table below) Satellite EGSE to be located in the LSV. [Note: Space limitations exist in the SEB, 350 ft umbilical cable length to spacecraft typical] Equipment Name / Type Approximate Size (LxWxH) Power Requirements

......

......

......

......

...... Is UPS required for equipment in the SEB? Yes / No Is Phone/Fax connection required in the SEB? Yes / No Circle: Phone / FAX LEV Describe (in the table below) Satellite EGSE to be located in the LEV. [Note: Space limitations exist in the SEB, 150 ft umbilical cable length to spacecraft typical] Equipment Name / Type Approximate Size (LxWxH) Power Requirements

......

......

......

......

...... Is UPS required for equipment in the LEV? Yes / No Is Ethernet connection between SEB and LEV required? Yes / No

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

SATELLITE SEPARATION REQUIREMENTS

ACCELERATION Longitudinal: = g’s Lateral: = g’s

VELOCITY Relative Separation Velocity Constraints:

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

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

± deg/sec Yaw: ± deg/sec ATTITUDE Describe Pointing Requirements Including Tolerances: (at deployment)

SPIN UP Longitudinal Spin Rate: ± deg/sec

OTHER Describe Any Other Separation Requirements:

SPACECRAFT COORDINATE SYSTEM

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

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SPACECRAFT PHYSICAL DIMENSIONS

STOWED Length/Height: Diameter: CONFIGURATION in cm in cm

Other Pertinent Dimension(s):

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

ON-ORBIT Describe size and shape: CONFIGURATION

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

SPACECRAFT MASS PROPERTIES*

2 2 PRE-SEPARATION Inertia units: lbm-in kg-m

Mass: lbm kg Ixx: Xcg: in cm Iyy: Izz: Ycg: in cm Ixy: Iyz: Zcg: in cm Ixz: 2 2 POST-SEPARATION Inertia units: lbm-in kg-m

(non-separating Mass: lbm kg adapter remaining with Ixx: launch vehicle) Xcg: in cm Iyy: Izz: Ycg: in cm Ixy: Iyz: Zcg: in cm Ixz: * Stowed configuration, spacecraft coordinate frame

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ASCENT TRAJECTORY REQUIREMENTS

Free Molecular Heating at Fairing Separation: Btu/ft2/hr (Standard Service: = 360 Btu/ft2/hr) FMH = W/m2 Fairing Internal Wall Temperature deg F (Standard Service: = 200°F) T = deg C 2 Dynamic Pressure at Fairing Separation: lbf /ft 2 2 (Standard Service: = 0.01 lbf /ft ) q = N/m 2 Ambient Pressure at Fairing Separation: lbf /in (Standard Service: = 0.3 psia) P = N/m2 2 Maximum Pressure Decay During Ascent: lbf /in /sec (Standard Service: = 0.6 psia) Δ P = N/m2/sec Thermal Maneuvers During Coast Periods: (Standard Service: none)

SPACECRAFT ENVIRONMENTS

THERMAL Spacecraft Thermal Dissipation, Pre-Launch Encapsulated: Watts DISSIPATION Approximate Location of Heat Source:

TEMPERATURE Temperature Limits During Max deg F deg C Ground/Launch Operations: Min deg F deg C (Standard Service is 55°F to 80°F) Component(s) Driving Temperature Constraint: Approximate Location(s):

HUMIDITY Relative Humidity: or, Dew Point: Max % Max deg F deg C Min % Min deg F deg C (Standard Service is 37 deg F) NITROGEN Specify Any Nitrogen Purge Requirements, Including Component Description, Location, PURGE and Required Flow Rate:

(Nitrogen Purge is a Non-Standard Service) CLEANLINESS Volumetric Requirements (e.g. Class 100,000): Surface Cleanliness (e.g. Visually Clean): Other: LOAD LIMITS Ground Transportation Load Limits: Axial = g’s Lateral = g’s

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ELECTRICAL INTERFACE

Bonding Requirements:

Are Launch Vehicle Supplied Pyro Commands Required? Yes / No If Yes, magnitude: amps for msec (Standard Service is 10 amps for 100 msec) Are Launch Vehicle Supplied If Yes, describe: Discrete Commands Required? Yes / No

Is Electrical Access to the Satellite Required... After Encapsulation? Yes / No at Launch Site Yes / No Is Satellite Battery Charging Required... After Encapsulation? Yes / No at Launch Site? Yes / No Is a Telemetry Interface with the Launch Vehicle Flight Computer Required? Yes / No

If Yes, describe:

Other Electrical Requirements:

Please complete attached sheet of required pass-through signals.

RF RADIATION

Time After Separation Until RF Devices Are Activated:

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

Frequency: MHz Power: Watts Location(s) on Satellite (spacecraft coordinate frame):

Longitudinal in cm Clocking (deg), Describe:

Longitudinal in cm Clocking (deg), Describe:

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REQUIRED PASS-THROUGH SIGNALS Max Total Line Item Current Resistance # Pin Signal Name From LEV To Satellite Shielding (amps) (ohms) 1

2

3

4

5

6

7

8

9

10

11

12

13

14

15

16

17

18

19

20

21

22

23

24

25

26

27

28

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MECHANICAL INTERFACE

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

SEPARATION Will Launch Vehicle Supply the Separation System? Yes / No SYSTEM If Yes approximate location of electrical connectors:

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

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

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

FAIRING Payload Fairing Access Doors (spacecraft coordinate frame): ACCESS Longitudinal in cm Clocking (deg), Describe:

Longitudinal in cm Clocking (deg), Describe:

Note: Standard Service is one door DYNAMICS Spacecraft Natural Frequency:

Axial Hz Lateral Hz

Recommended: > TBD Hz > TBD Hz OTHER Other Mechanical Interface Requirements:

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