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March 2002 Release 1.0

Approved for Public Release Distribution Unlimited

March 2002 Minotaur® Users Guide Release 1.0

Approved for Public Release Distribution Unlimited

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

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

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

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

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

Business Development Orbital Sciences Corporation Launch Systems Group 3380 S. Price Road Chandler, AZ 85248

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

Release 1.0 March 2002 Minotaur User's Guide

TABLE OF CONTENTS

SECTION PAGE 1.0 INTRODUCTION ...... 1-1

2.0 MINOTAUR LAUNCH SERVICE...... 2-1 2.1 Minotaur Launch System Overview ...... 2-1 2.2 Minotaur Launch Service ...... 2-1 2.3 Minotaur Launch Vehicle ...... 2-2 2.3.1 Lower Stack Assembly ...... 2-2 2.3.2 Upper Stack Assembly ...... 2-3 2.4 Launch Support Equipment ...... 2-4

3.0 GENERAL PERFORMANCE ...... 3-1 3.1 Mission Profiles ...... 3-1 3.2 Launch Sites...... 3-1 3.2.1 Western Launch Sites ...... 3-1 3.2.2 Eastern Launch Sites ...... 3-2 3.3 Performance Capability...... 3-2 3.4 Injection Accuracy ...... 3-2 3.5 Payload Deployment ...... 3-5 3.6 Payload Separation ...... 3-5 3.7 Collision/Contamination Avoidance Maneuver ...... 3-6

4.0 PAYLOAD ENVIRONMENT ...... 4-1 4.1 Steady State and Transient Acceleration Loads ...... 4-1 4.1.1 Optional Payload Isolation System...... 4-2 4.2 Payload Vibration Environment ...... 4-3 4.3 Payload Shock Environment ...... 4-3 4.4 Payload Acoustic Environment ...... 4-3 4.5 Payload Structural Integrity and Environments Verification ...... 4-4 4.5.1 Recommended Payload Testing and Analysis ...... 4-4 4.6 Thermal and Humidity Environments ...... 4-5 4.6.1 Ground Operations ...... 4-5 4.6.2 Powered Flight ...... 4-6 4.6.3 Nitrogen Purge (non-standard service) ...... 4-7 4.7 Payload Contamination Control...... 4-7 4.8 Payload Electromagnetic Environment ...... 4-7

5.0 PAYLOAD INTERFACES ...... 5-1 5.1 Payload Fairing ...... 5-1 5.1.1 Payload Dynamic Design Envelope ...... 5-1 5.1.2 Payload Access Door ...... 5-1 5.1.3 Increased Volume Payload Fairing ...... 5-1 5.2 Payload Mechanical Interface and Separation System ...... 5-1

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TABLE OF CONTENTS

SECTION PAGE

5.2.1 Standard Non-Separating Mechanical Interface...... 5-1 5.2.2 Separating Mechanical Interface ...... 5-5 5.2.3 Orbital Supplied Drill Templates ...... 5-5 5.3 Payload Electrical Interfaces...... 5-5 5.3.1 Payload Umbilical Interfaces...... 5-5 5.3.2 Payload Battery Charging ...... 5-5 5.3.3 Payload Command and Control ...... 5-5 5.3.4 Pyrotechnic Initiation Signals...... 5-10 5.3.5 Payload Telemetry ...... 5-10 5.3.6 Non Standard Electrical Interfaces ...... 5-10 5.3.7 Electrical Launch Support Equipment ...... 5-10 5.4 Payload Design Constraints ...... 5-10 5.4.1 Payload Center of Mass Constraints ...... 5-10 5.4.2 Final Mass Properties Accuracy ...... 5-10 5.4.3 Pre-Launch Electrical Constraints ...... 5-10 5.4.4 Payload EMI/EMC Constraints...... 5-10 5.4.5 Payload Dynamic Frequencies...... 5-11 5.4.6 Payload Propellent Slosh...... 5-11 5.4.7 Payload-Supplied Separation Systems ...... 5-11 5.4.8 System Safety Constraints ...... 5-11

6.0 MISSION INTEGRATION ...... 6-1 6.1 Mission Management Approach ...... 6-1 6.1.1 SMC Det 12/RP Mission Responsibilities ...... 6-1 6.1.2 Orbital Mission Responsibilities ...... 6-1 6.2 Mission Planning and Development ...... 6-2 6.3 Mission Integration Process ...... 6-2 6.3.1 Integration Meetings ...... 6-2 6.3.2 Mission Design Reviews ...... 6-4 6.3.3 Readiness Reviews ...... 6-4 6.4 Documentation ...... 6-4 6.4.1 Customer-Provided Documentation ...... 6-4 6.4.1.1 Payload Questionnaire ...... 6-4 6.4.1.2 Payload Mass Properties ...... 6-4 6.4.1.3 Payload Finite Element Model ...... 6-5 6.4.1.4 Payload Thermal Model for Integrated Thermal Analysis...... 6-5 6.4.1.5 Payload Drawings...... 6-5 6.4.1.6 Program Requirements Document (PRD) Mission Specific Annex Inputs ...... 6-5 6.4.1.6.1 Launch Operations Requirements (OR) Inputs ...... 6-5 6.5 Safety ...... 6-5 6.5.1 System Safety Requirements ...... 6-5 6.5.2 System Safety Documentation...... 6-6

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SECTION PAGE

7.0 GROUND AND LAUNCH OPERATIONS ...... 7-1 7.1 Minotaur/Payload Integration Overview ...... 7-1 7.2 Ground And Launch Operations ...... 7-1 7.2.1 Launch Vehicle Integration ...... 7-1 7.2.1.1 Planning and Documentation ...... 7-1 7.2.1.2 Vehicle Integration and Test Activities ...... 7-1 7.2.1.2.1 Flight Simulation Tests ...... 7-1 7.2.2 Payload Processing/Integration ...... 7-1 7.2.2.1 Payload to Minotaur Integration ...... 7-2 7.2.2.2 Pre-Mate Interface Testing ...... 7-2 7.2.2.3 Payload Mating and Verification ...... 7-2 7.2.2.4 Final Processing and Fairing Closeout ...... 7-2 7.2.2.5 Payload Propellant Loading ...... 7-2 7.2.2.6 Final Vehicle Integration and Test ...... 7-2 7.3 Launch Operations ...... 7-2 7.3.1 Launch Control Organization ...... 7-2

8.0 OPTIONAL ENHANCED CAPABILITIES ...... 8-1 8.1 Mechanical Interface and Separation System Enhancements ...... 8-1 8.1.1 Separation Systems ...... 8-1 8.1.2 Additional Fairing Access Doors ...... 8-1 8.1.3 Payload Isolation System ...... 8-1 8.1.4 Increased Payload Volume...... 8-1 8.2 Performance Enhancements ...... 8-2 8.2.1 Insertion Accuracy ...... 8-2 8.3 Environmental Control Options ...... 8-2 8.3.1 Conditioned Air ...... 8-2 8.3.2 Nitrogen Purge ...... 8-2 8.3.3 Enhanced Contamination Control ...... 8-3 8.3.3.1 High Cleanliness Integration Environment (Class 10K or 100K) ...... 8-3 8.3.3.2 Fairing Surface Cleanliness Options ...... 8-3 8.3.3.3 High Cleanliness Fairing Environment ...... 8-3 8.4 Enhanced Telemetry Options ...... 8-3 8.4.1 Enhanced Telemetry Bandwidth ...... 8-3 8.4.2 Enhanced Telemetry Instrumentation ...... 8-3 8.4.3 Navigation Data ...... 8-4

9.0 SHARED LAUNCH ACCOMMODATIONS ...... 9-1 9.1 Load-Bearing Spacecraft ...... 9-1 9.2 Non Load-Bearing Spacecraft...... 9-2

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SECTION PAGE

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

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

LIST OF FIGURES

FIGURE PAGE

Figure 2-1 OSP Minotaur Launch Vehicle ...... 2-1 Figure 2-2 OSP Minotaur Launch Vehicle Configuration ...... 2-2 Figure 2-3 Minotaur Upper Stack Assembly Processing at Orbital's Vehicle Assembly Building at VAFB ...... 2-3 Figure 2-4 Minotaur EGSE Configuration ...... 2-5 Figure 2-5 Minotaur Launch Control Consoles Configuration ...... 2-5 Figure 3-1 Minotaur Typical Mission Profile ...... 3-1 Figure 3-2 Minotaur Launch Site Options ...... 3-2 Figure 3-3 Minotaur Performance - California Spaceport Launches (SSI CLF) ...... 3-3 Figure 3-4 Minotaur Performance - Kodiak Launch Complex Launches ...... 3-3 Figure 3-5 Minotaur Performance - Spaceport Florida Launches ...... 3-4 Figure 3-6 Minotaur Performance - Virginia Spaceflight Center Launches ...... 3-4 Figure 3-7 Stage Impact Points for Typical Sun-Synchronous Launch From VAFB ...... 3-5 Figure 3-8 Injection Accuracies to Low Earth Orbits ...... 3-5 Figure 3-9 Typical Pre-Separation Payload Pointing and Spin Rate Accuracies ...... 3-5 Figure 4-1 Phasing of Dynamic Loading Events ...... 4-1 Figure 4-2 Payload Design CG Net Load Factors (Typical)...... 4-1 Figure 4-3 Minotaur 3-Sigma High Maximum Acceleration as a Function of Payload Weight ...... 4-2 Figure 4-4 Payload Random Vibration Environment During Flight ...... 4-3 Figure 4-5 Maximum Shock Environment - Launch Vehicle to Payload...... 4-3 Figure 4-6 Maximum Shock Environment - Payload to Launch Vehicle...... 4-3 Figure 4-7 Payload Acoustic Environment During Liftoff and Flight ...... 4-4 Figure 4-8 Factors of Safety Payload Design and Test ...... 4-4 Figure 4-9 Recommended Payload Testing Requirements...... 4-5 Figure 4-10 Payload Thermal and Humidity Environments ...... 4-5 Figure 4-11 Minotaur Worst-Case Payload Fairing Inner Surface Temperature During Ascent (Payload Region) ...... 4-6 Figure 4-12 Minotaur Launch Vehicle RF Emitters and Receivers ...... 4-8

Release 1.0 March 2002 iv Minotaur User's Guide

LIST OF FIGURES CONTINUED

FIGURE PAGE Figure 5-1 Payload Fairing Dynamic Envelope with 38 in (97 cm) Diameter Payload Interface ...... 5-2 Figure 5-2 Payload Fairing Access Door Placement Zone ...... 5-3 Figure 5-3 Non-Separable Payload Mechanical Interface ...... 5-4 Figure 5-4 38 in (97 cm) Separable Payload Interface ...... 5-6 Figure 5-5 23 in (59 cm) Separable Payload Interface ...... 5-7 Figure 5-6 17 in (43 cm) Separable Payload Interface ...... 5-8 Figure 5-7 Payload Separation Velocities Using the Standard Separation System ...... 5-9 Figure 5-8 Vehicle/Spacecraft Electrical Connectors and Associated Electrical Harnesses...... 5-9 Figure 5-9 Payload Mass Properties Measurement Tolerance ...... 5-10 Figure 6-1 OSP Management Structure...... 6-1 Figure 6-2 Typical Minotaur Mission Integration Schedule ...... 6-3 Figure 8-1 Softride for Small Satellites (SRSS) Payload Isolation System ...... 8-1 Figure 8-2 Optional 61 in. Diameter Fairing ...... 8-2 Figure 9-1 Typical Load Bearing Spacecraft Configuration ...... 9-1 Figure 9-2 JAWSAT Multiple Payload Adapter Load Bearing Spacecraft ...... 9-2 Figure 9-3 Five Bay Multiple Payload Adapter Concept ...... 9-2 Figure 9-4 Dual Payload Attach Fitting Configuration ...... 9-3 Figure B-1 Typical Minotaur Payload Electrical Interface Block Diagram ...... B-2

Release 1.0 March 2002 v Minotaur Payload User's Guide Glossary

A Ampere g Gravitational Force AADC Alaska Aerospace Development GACS Ground Air Conditioning System Corporation GFE Government Furnished Equipment AC Air Conditioning GN2 Gaseous Nitrogen ACS Attitude Control System GPB GPS Position Beacon AFRL Air Force Research Laboratory GPS Global Positioning System AODS All-Ordnance Destruct System GSE Ground Support Equipment ATP Authority to Proceed GTO Geosynchronous Transfer Orbit C/CAM Collision/Contamination Avoidance HAPS Hydrazine Auxiliary Propulsion System Maneuver HEPA High Efficiency Particulate Air CCAS Cape Canaveral Air Station Hz Hertz CDR Critical Design Review I/F Interface CFE Customer Furnished Equipment ICD Interface Control Drawing CG,cg Center of Gravity ILC Initial Launch Capability CLA Coupled Loads Analysis IMU Inertial Measurement Unit CLF Commercial Launch Facility in Inch cm Centimeter INS Inertial Navigation System CVCM Collected Volatile Condensable Materials ISO International Standardization Organization dB Decibels IVT Interface Verification Test deg Degrees kbps Kilobits per Second DPAF Dual Payload Attach Fitting kg Kilograms ECS Environmental Control System km Kilometer EGSE Electrical Ground Support Equipment lb Pound EMC Electromagnetic Compatibility lbm Pound(s) of Mass EME Electromagnetic Environment LCR Launch Control Room EMI Electromagnetic Interference LEV Launch Equipment Vault FLSA Florida Spaceport Authority LITVC Liquid Injection Thrust Vector Control FM Frequency Modulation LOCC Launch Operations Control Center ft Feet LRR Launch Readiness Review FTLU Flight Termination Logic Unit LSA Lower Stack Assembly FTS Flight Termination System LSE Launch Support Equipment

Release 1.0 March 2002 vi Minotaur Payload User's Guide Glossary m/s Meters per Second POC Point of Contact

Mbps Mega Bits per Second ppm Parts Per Million mA Milliamps PRD Program Requirements Document MACH Modular Avionics Control Hardware PSD Power Spectral Density MDR Mission Design Review PSP Prelaunch Safety Package MHz MegaHertz RCS Roll Control System MIL-STD Military Standard RF Radio Frequency MIWG Mission Integration Working Group mm Millimeter RGIU Rate Gyro Interface Unit

MPA Multiple Payload Adapter RGU Rate Gyro Unit

MRD Mission Requirements Document rpm Revolutions per Minute

MRR Mission Readiness Review RSLP Rocket Systems Launch Program

MSPSP Missile System Prelaunch Safety RWG Range Working Group Package ms Millisecond s&a Safe & Arm

N/A Not Applicable scfm Standard Cubic Feet per Minute

NCU Nozzle Control Unit SEB Support Equipment Building

NM Nautical Mile sec Second(s)

OD Operations Directive SINDA Finite Element Thermal Analysis Tool OD Outside Dimension Tradename

ODM Ordnance Driver Module SLC Space Launch Complex

OR Operations Requirements Document SLV Space Launch Vehicle

OSP Orbital Suborbital Program SMC Space and Missile Systems Center

P/L Payload SOC Statement of Capabilities

PACS Pad Air Conditioning System SPL Sound Pressure Level PAF Payload Attach Fitting SRM Solid Rocket Motor PCM Pulse Code Modulation SRS Shock Response Spectrum PDR Preliminary Design Review PI Program Introduction SRSS Softride for Small Satellites PID Proportional-Integral-Derivative SSI Spaceport Systems International

Release 1.0 March 2002 vii Minotaur Payload User's Guide Glossary

STA Station TLV Target Launch Vehicle TML Total Mass Loss TVC Thrust Vector Control UDS Universal Documentation System UFS Ultimate Factory of Safety USAF United States Air Force V/M Volts per Meter VAB Vehicle Assembly Building VAFB Vandenberg Air Force Base W Watt WFF Wallops Flight Facility WP Work Package YFS Yield Factor of Safety

Release 1.0 March 2002 viii Minotaur Payload User's Guide Section 1.0 - Introduction

1.0 INTRODUCTION User’s Guide describes Minotaur-unique This User’s Guide is intended to integration and test approaches (including the familiarize payload mission planners with the typical operational timeline for payload capabilities of the Orbital Suborbital Program integration with the Minotaur vehicle) and the (OSP) Minotaur Space Launch Vehicle (SLV) existing ground support equipment that is used to launch service. This document provides an conduct Minotaur operations. overview of the Minotaur system design and a description of the services provided to our customers. Minotaur offers a variety of enhanced options to allow the maximum flexibility in satisfying the objectives of single or multiple payloads.

The Minotaur’s primary mission is to provide low cost, high reliability launch services to government-sponsored payloads. Minotaur accomplishes this using flight-proven components with a significant flight heritage such as surplus Minuteman II boosters, the upper stage Pegasus motors, the Pegasus Fairing and Attitude Control System, and a mix of Pegasus, Taurus, and suborbital 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 Minotaur missions to date.

The Minotaur launch vehicle system is composed of a flight vehicle and ground support equipment. Each element of the Minotaur 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 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 mission integration process.

The Minotaur system can operate from a wide range of launch facilities and geographic locations. The system is compatible with, and will typically operate from, commercial spaceport facilities and existing U.S. Government ranges at Vandenberg Air Force Base (VAFB), Cape Canaveral Air Station (CCAS), Wallops Flight Facility (WFF), and Kodiak Island, Alaska. This

Release 1.0 March 2002 1-1 Minotaur Payload User's Guide Section 2.0 - Minotaur Launch Service

2. MINOTAUR LAUNCH SERVICE transportability, and operation from multiple launch sites. Minotaur draws on the successful 2.1. Minotaur Launch System Overview heritage of three launch vehicles: Orbital’s The Minotaur launch vehicle, shown in Pegasus and Taurus space launch vehicles and Figure 2-1, was developed by Orbital for the the Minuteman II system of the USAF. Minotaur’s United States Air Force (USAF) to provide a cost upper two stages and avionics are derived from effective, reliable and flexible means of placing the Pegasus and Taurus systems, providing a small satellites into orbit. An overview of the combined total of more than 30 successful space system and available launch services is provided launch missions. Orbital’s efforts have enhanced within this section, with specific elements covered or updated Pegasus and Taurus avionics in greater detail in the subsequent sections of this components to meet the payload-support User’s Guide. requirements of the OSP program. Combining these improved subsystems with the long Minotaur has been designed to meet the successful history of the Minuteman II boosters needs of United States Government-sponsored has resulted in a simple, robust, self-contained customers at a lower cost than commercially launch system that has been completely available alternatives by the use of surplus successful in its flights to date and is fully Minuteman boosters. The requirements of that operational to support government-sponsored program stressed system reliability, small satellite launches.

The Minotaur system also includes a complete set of transportable Launch Support Equipment (LSE) designed to allow Minotaur 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 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. This has been the case for the first launches from VAFB.

The vehicle and LSE are designed to be capable of launch from any of the four commercial Spaceports (Alaska, California, Florida, and Virginia), as well as from existing U.S. Government facilities at VAFB and CCAS. The Launch Control Room (LCR) serves as the actual control center for conducting a Minotaur launch and includes consoles for Orbital, range safety, and customer personnel. Further description of the Launch Support Equipment is provided in Section 2.4.

2.2. Minotaur Launch Service Figure 2-1. OSP Minotaur Launch Vehicle The Minotaur Launch Service is provided

Release 1.0 March 2002 2-1 Minotaur Payload User's Guide Section 2.0 - Minotaur Launch Service through the combined efforts of the USAF and Minotaur mission integration process completely Orbital, along with associate contractors including identifies, documents, and verifies all spacecraft TRW and Commercial Spaceports. The primary and mission requirements. This provides a solid customer interface will be with the USAF Space basis for initiating and streamlining the integration and Missile Systems Center, Detachment 12, process for future Minotaur customers. Rocket Systems Launch Program (RSLP), designated SMC Det 12/RP. Orbital is the launch 2.3. Minotaur Launch Vehicle vehicle provider. For brevity, this integrated The Minotaur vehicle, shown in expanded team effort will be referred to as “OSP”. Where view in Figure 2-2, is a four stage, inertially interfaces are directed toward a particular member guided, all solid propellant ground launched of the team, they will be referred to directly (i.e. vehicle. Conservative design margins, state-of- “Orbital” or “SMC Det 12/RP”). the-art structural systems, a modular avionics architecture, and simplified integration and test OSP provides all of the necessary capability, yield a robust, highly reliable launch hardware, software and services to integrate, test vehicle design. In addition, Minotaur payload and launch a satellite into its prescribed orbit. In accommodations and interfaces have been addition, OSP will complete all the required designed to satisfy a wide range of potential agreements, licenses and documentation to payload requirements. successfully conduct Minotaur operations. All Minotaur production and integration processes 2.3.1. Lower Stack Assembly and procedures have been demonstrated and are The Lower Stack Assembly (LSA) consists in place for future Minotaur missions. The of the refurbished Government Furnished

Figure 2-2. OSP Minotaur Launch Vehicle Configuration

Release 1.0 March 2002 2-2 Minotaur Payload User's Guide Section 2.0 - Minotaur Launch Service

Equipment (GFE) Minuteman Stages 1 and 2. subsystems, the LSE, and the payload, if required, Only minor modifications are made to the utilizing standard RS-422 serial links and discrete boosters, including harness interface changes. I/O. The Minotaur design incorporates Orbital’s Modular Avionics Control Hardware (MACH) to The first stage consists of the Minuteman provide power transfer, data acquisition, II M55A1 solid propellant motor, Nozzle Control Minuteman booster interfaces, and ordnance Units (NCU), Stage 1 Ignition Safe/Arm, S1/S2 initiation. MACH has exhibited 100% reliability Interstage and Stage 1 FTS. Four gimbaled nozzles on OSP SLV and Target Launch Vehicle (TLV) provide three axis control during first stage burn. flights and several of Orbital’s suborbital launch The Second stage consists of a refurbished vehicles. In addition, the Minotaur telemetry Minuteman II SR19 motor, Liquid Injection Thrust system has been upgraded to provide up to 2 Vector Control subsystem (LITVC), S2 ignition Mbps of real-time vehicle data with dedicated safe/arm device, a Roll Control System (RCS), and bandwidth and channels reserved for payload the Stage 2 FTS components. Attitude control use. during second stage burn is provided by the operational LITVC and hot gas roll control. A Attitude Control Systems –– The Minotaur Rate Gyro Unit (RGU) is installed on the outer Attitude Control System (ACS) provides three- skin of the SR19 to enhance the vehicle control axis attitude control throughout boosted flight and increase launch availability. and coast phases. Stages 1 and 2 utilize the Minuteman Thrust Vector Control (TVC) systems. 2.3.2. Upper Stack Assembly The Stage 1 TVC is a four-nozzle hydraulic system, The Minotaur Upper Stack is composed while the Stage 2 system combines liquid of Stages 3 and 4 which are the Alliant injection for pitch and yaw control with hot gas TechSystems Orion 50XL and 38 SRMs, roll control. Stages 3 and 4 utilize the same TVC respectively. These motors were originally systems as Pegasus and Taurus. They combine developed for Orbital’s Pegasus program and single-nozzle electromechanical TVC for pitch have been adapted for use on the ground- and yaw control with a three-axis cold-gas launched Minotaur vehicle. Common design attitude control system resident in the avionics features, materials and production techniques section providing roll control. are applied to both motors to maximize reliability and production efficiency. The motors are fully Attitude control is achieved using a three- flight qualified based on their heritage, design axis autopilot that employs Proportional-Integral- conservatism, ground static fires and over thirty successful flights. Processing of the Minotaur Upper Stack is conducted at the same processing facility as Pegasus, directly applying the integration and testing experience of Pegasus to the Minotaur system (Figure 2-3).

Avionics — The Minotaur avionics system has heritage to the Pegasus and Taurus designs. However, the Minotaur design was upgraded to provide the increased capability and flexibility required by the OSP contract, particularly in the area of payload accommodations. The flight computer, which is common to Pegasus and Figure 2-3. Minotaur Upper Stack Assembly Taurus, is a 32-bit multiprocessor architecture. It Processing at Orbital's Vehicle Assembly provides communication with vehicle Building at VAFB

Release 1.0 March 2002 2-3 Minotaur Payload User's Guide Section 2.0 - Minotaur Launch Service

Derivative (PID) control. Stages 1 and 2 fly a pre- With the addition of a structural adapter, programmed attitude profile based on trajectory either fairing can accommodate multiple design and optimization. Stage 3 uses a set of pre- payloads. This feature, described in more detail programmed orbital parameters to place the vehicle in Section 9.0 of this User’s Guide, permits two or on a trajectory toward the intended insertion apse. more smaller payloads to share the cost of a The extended coast between Stages 3 and 4 is used Minotaur launch, resulting in a lower launch cost to orient the vehicle to the appropriate attitude for for each as compared to other launch options. Stage 4 ignition based upon a set of pre-programmed 2.4. Launch Support Equipment orbital parameters and the measured performance The Minotaur LSE is designed to be of the first three stages. Stage 4 utilizes energy readily adaptable to varying launch site management to place the vehicle into the proper configurations with minimal unique orbit. After the final boost phase, the three-axis cold- infrastructure required. The EGSE consists of gas attitude control system is used to orient the readily transportable consoles that can be vehicle for spacecraft separation, contamination housed in various facility configurations and collision avoidance and downrange downlink depending on the launch site infrastructure. The maneuvers. The autopilot design is modular, so EGSE is composed of two primary functional additional payload requirements such as rate control elements: Launch Control and Vehicle Interface or celestial pointing can be accommodated with (Figure 2-4). The Launch Control consoles are minimal additional development. located in a LCR, depending on available launch site accommodations. The Vehicle Interface Telemetry Subsystem –– The Minotaur EGSE is located in structures near the pad, typically called a Support Equipment Building telemetry subsystem provides real-time health (SEB). Fiber optic connections from the Launch and status data of the vehicle avionics system, as Control to the Vehicle Interface consoles are well as key information regarding the position, used for efficient, high bandwidth performance and environment of the Minotaur communications and to minimize the amount of vehicle. This data may be used by Orbital and the cabling required. The Vehicle Interface racks range safety personnel to evaluate system provide the junction from fiber optic cables to performance. The minimum data rate is 750 kbps, the copper cabling interfacing with the vehicle. but the system is capable of data rates up to 2 Mbps. The LCR serves as the control center Payload Fairings –– The baseline Minotaur during the launch countdown. The number of fairing is identical to the Pegasus fairing design. personnel that can be accommodated are However, due to differences in vehicle loads and dependent on the launch site facilities. At a environments, the Minotaur implementation minimum, the LCR will accommodate Orbital allows a larger payload envelope. The Minotaur personnel controlling the vehicle, two Range payload fairing consists of two composite shell Safety representatives (ground and flight safety), halves, a nose cap integral to a shell half, and a and the Air Force Mission Manager. A typical layout is shown in Figure 2-5. Mission-unique, separation system. Each shell half is composed of customer-supplied payload consoles and a cylinder and ogive sections. equipment can be supported in the LCR and SEB, within the constraints of the launch site Options for payload access doors and facilities or temporary structure facilities. enhanced cleanliness are available. A larger 61 Interface to the payload through the Minotaur inch diameter (OD) fairing is also in development payload umbilicals and land lines provides the and is available for future missions. Further details capability for direct monitoring of payload on the baseline fairing are included in Section 5.1 functions. Payload personnel accommodations and for the larger fairing in Section 8.1. will be handled on a mission-specific basis.

Release 1.0 March 2002 2-4 Minotaur Payload User's Guide Section 2.0 - Minotaur Launch Service

Launch Control Room

Telemetry Monitor Console

Data Reduction Console Support Equipment Building Background Limit Checking Console

Flight Computer Console Power Supply Rack

Fiber Data Display Copper Lines Console Serial FTS Interface Interface Patch Panel Rack Rack

Power Control Console Arming/ Ignition Rack Data Display Console Payload I/F Rack FTS Control Console

Payload TM14025_059 Console

Figure 2-4. Minotaur EGSE Configuration

Figure 2-5. Minotaur Launch Control Consoles Configuration

Release 1.0 March 2002 2-5 Minotaur Payload User's Guide Section 3.0 - General Performance

3. GENERAL PERFORMANCE 3.2. Launch Sites Depending on the specific mission and 3.1. Mission Profiles range safety requirements, Minotaur can operate Minotaur can attain a range of posigrade from several East and West launch sites, illustrated and retrograde inclinations through the choice of in Figure 3-2. Specific performance parameters launch sites made available by the readily are presented in Section 4. adaptable nature of the Minotaur launch system. A typical mission profile to a sun-synchronous 3.2.1. Western Launch Sites orbit is shown in Figure 3-1. High energy and For missions requiring high inclination Geosynchronous Transfer Orbit (GTO) missions orbits (greater than 60°), launches can be can also be achieved. All performance parameters presented herein are typical for most expected conducted from facilities at VAFB, CA, or Kodiak payloads. However, performance may vary Island, AK. Both facilities can accommodate ° ° depending on unique payload or mission inclinations from 60 to 120 , although ° characteristics. Specific requirements for a inclinations below 72 from VAFB would require particular mission should be coordinated with an out-of-plane dogleg, thereby reducing payload OSP. Once a mission is formally initiated, the capability. Initial Minotaur missions were requirements will be documented in the Mission launched from the California Spaceport facility, Requirements Document (MRD). Further detail operated by Spaceport Systems International will be captured in the Payload-to-Launch Vehicle (SSI), on South VAFB, near SLC-6. The launch Interface Control Drawing (ICD). facility at Kodiak Island, operated by the Alaska

Figure 3-1. Minotaur Typical Mission Profile

Release 1.0 March 2002 3-1 Minotaur Payload User's Guide Section 3.0 - General Performance

KODIAK LAUNCH COMPLEX VIRGINIA SPACE FLIGHT Kodiak Island, AK CENTER Wallops Island, VA • Commercial Launch Sites at NASA's Wallops Flight Facilities

WESTERN RANGE EASTERN RANGE Vandenberg AFB, CA Patrick AFB, FL • Government Launch Sites • Government Launch Sites • California Spaceport SSI CLF • Spaceport Florida TM14025_081 Figure 3-2. Minotaur Launch Site Options

Aerospace Development Corporation (AADC) 3.3. Performance Capability has been used for both orbital and suborbital Minotaur performance curves for circular launches. and elliptical orbits of various altitudes and inclinations are detailed in Figure 3-3 through 3.2.2. Eastern Launch Sites Figure 3-6 for launches from all four Spaceports. For Easterly launch azimuths to achieve These performance curves provide the total mass above the standard, non-separating orbital inclinations between 28.5° and 60°, interface. The mass of any Payload Attach Fitting Minotaur can be launched from facilities at Cape (PAF) or separation system is to be accounted for Canaveral, FL or Wallops Island, VA. Launches in the payload mass allocation. Figure 3-7 from Florida will nominally use the Florida illustrates the stage vacuum impact points for a Spaceport Authority (FLSA) launch facilities at typical sun-synchronous trajectory from VAFB. LC-46 on CCAS, Cape Canaveral, FL. These will be typically for inclinations from 28.5° to 40°, 3.4. Injection Accuracy although inclinations above 35° may have Minotaur injection accuracy is summarized in Figure 3-8. Better accuracy can reduced performance due to the need for a be provided dependent on specific mission trajectory dogleg. The Virginia Spaceflight Center characteristics. For example, heavier payloads facilities at the WFF may be used for inclinations will typically have better insertion accuracy, as ° ° from 30 to 60 . Southeasterly launches from will higher orbits. An enhanced option for WFF offer fewer overflight concerns than CCAS. increased insertion accuracy is also available Inclinations below 35° and above 55° are feasible, (Section 8.2.1). It utilizes the flight-proven albeit with doglegs and altitude constraints due Hydrazine Auxiliary Propulsion System (HAPS) to stage impact considerations. developed on the Pegasus program.

Release 1.0 March 2002 3-2 Minotaur Payload User's Guide Section 3.0 - General Performance

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Figure 3-5. Minotaur Performance - Spaceport Florida Launches

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

Release 1.0 March 2002 3-4 Minotaur Payload User's Guide Section 3.0 - General Performance

Stage 1 30

Stage 2 20 Latitude (deg) 10

10 Stage 3 20 180 170 160 150 140 130 120 110 100 Longitude (deg) TM14025_066

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

Tolerance 3.5. Payload Deployment Error Type (Worst Case) Following orbit insertion, the Minotaur Stage 4 avionics subsystem can execute a series of ACS maneuvers to provide the desired initial Altitude (Insertion Apse) ±10 nm (18.5 km) payload attitude prior to separation. This Altitude (Non-Insertion ±50 nm (92.6 km) capability may also be used to incrementally Apse) reorient Stage 4 for the deployment of multiple spacecraft with independent attitude Altitude (Mean) ± 30 nm (55.6 km) requirements. Either an inertially-fixed or spin- Inclination ±0.2° stabilized attitude may be specified by the customer. Figure 3-8. Injection Accuracies to Low Earth Orbits The maximum spin rate for a specific mission depends upon the spin axis moment of Error inertia of the payload and the amount of ACS Type Angle Rate propellant needed for other attitude maneuvers. 3-Axis Yaw ±1.0° ±0.5°/sec Figure 3-9 provides the typical payload pointing and spin rate accuracies. Pitch ±1.0° ±0.5°/sec Roll ±1.0° ±0.5°/sec 3.6. Payload Separation Spinning Spin Axis ±1.0° ≤10 rpm Payload separation dynamics are highly dependent on the mass properties of the payload Spin Rate 3°/sec and the particular separation system utilized. Figure 3-9. Typical Pre-Separation Payload The primary parameters to be considered are Pointing and Spin Rate Accuracies payload tip-off and the overall separation velocity.

Release 1.0 March 2002 3-5 Minotaur Payload User's Guide Section 3.0 - General Performance

Payload tip-off refers to the angular velocity imparted to the payload upon separation due to payload 8 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 4°/sec per axis. Orbital performs a mission-specific tip-off analysis for each payload.

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

3.7. Collision/Contamination Avoidance Maneuver Following orbit insertion and payload separation, the Minotaur Stage 4 will perform a Collision/Contamination Avoidance Maneuver (C/CAM). The C/CAM minimizes both payload contamination and the potential for recontact between Minotaur hardware and the separated payload. OSP will perform a recontact analysis for post separation events.

A typical C/CAM begins soon after payload separation. The launch vehicle performs a 90° yaw maneuver designed to direct any remaining Stage 4 motor impulse in a direction which will increase the separation distance between the two bodies. After a delay to allow the distance between the spacecraft and Stage 4 to increase to a safe level, the launch vehicle begins a “crab-walk” maneuver to impart a small amount of delta velocity, increasing the separation between the payload and the fourth stage of the Minotaur.

Following the completion of the C/CAM maneuver as described above and any remaining maneuvers, such as downlinking of delayed telemetry data, the ACS valves are opened and the remaining ACS nitrogen propellent is expelled.

Release 1.0 March 2002 3-6 Minotaur Payload User's Guide Section 4.0 - Payload Environment

4. PAYLOAD ENVIRONMENT Dynamic loading events that occur This section provides details of the pre- throughout various portions of the flight in- dicted environmental conditions that the pay- clude steady state acceleration, transient low load will experience during Minotaur ground frequency acceleration, acoustic impingement, operations, powered flight, and launch system random vibration, and pyrotechnic shock on-orbit operations. events. Figure 4-1 identifies the time phasing of these dynamic loading events and environ- Minotaur ground operations include ments and their significance. Pyroshock events payload integration and encapsulation within are not indicated in this figure, as they do not the fairing, subsequent transportation to the occur simultaneous with any other significant launch site and final vehicle integration activi- dynamic loading events. ties. Powered flight begins at Stage 1 ignition and ends at Stage 4 burnout. Minotaur on-or- 4.1. Steady State and Transient Acceleration bit operations begin after Stage 4 burnout and Loads end following payload separation. To more Design limit load factors due to the accurately define simultaneous loading and combined effects of steady state and low environmental conditions, the powered flight frequency transient accelerations are defined portion of the mission is further subdivided into in Figure 4-2. These values include uncertainty smaller time segments bounded by critical margins and are typical for an 800 lbm flight events such as motor ignition, stage sepa- payload. ration, and transonic crossover. Event Axial (G's) Lateral (G's) The environmental design and test crite- Liftoff 0 ±4.6 0 ±1.6 ria presented have been derived using measured Transonic 3.2 ±0.5 0.2 ±1.2 data obtained from previous Pegasus, Taurus and Minotaur missions, motor static fire tests, other Supersonic 3.8 ±0.5 0.4 ±1.1 system development tests and analyses. These S2 Ignition 0 ±6.6 0 ±3.3 criteria are applicable to Minotaur configurations S3 Ignition 0 ±6.1 0 ±0.5 using the standard 50 in. diameter fairing. The predicted levels presented are intended to be rep- S3 Burnout See Fig 4-3 0.3 ±0.0 resentative of mission specific levels. Mission S4 Burnout See Fig 4-3 0.3 ±0.0 specific analyses that are performed as a stan- TM14025_068 dard service are documented or referenced in Figure 4-2. Payload Design CG Net Load the mission ICD. Factors (Typical)

Supersonic/ Balance of S2 S2 S3 S3 S4 Item Liftoff Transonic Max Q S1 Burn Ignition Burn Ignition Burn Burn

Typical Flight 3 sec 17 sec 30 sec 10 sec N/A 70 sec N/A 70 sec 70 sec Duration Steady State Loads Yes Yes Yes Yes No Yes No Yes Yes Transient Loads Yes Yes Yes No Yes No Yes No No Acoustics Yes Yes Yes Yes No No No No No Random Vibration Yes Yes Yes Yes No Yes No Yes Yes

TM14025_067 Figure 4-1. Phasing of Dynamic Loading Events

Release 1.0 March 2002 4-1 Minotaur Payload User's Guide Section 4.0 - Payload Environment

During powered flight, the maximum liminary design purposes, Orbital can provide steady state accelerations are dependent on the preliminary Center-of-Gravity (CG) netloads payload mass. The maximum level can po- given a payload’s mass properties, CG loca- tentially occur in either Stage 3 or 4 burn. Fig- tion and bending frequencies. ure 4-3 depicts the axial acceleration at burnout for each stage as a function of payload 4.1.1. Optional Payload Isolation System mass. OSP offers a flight-proven payload load isolation system as a non-standard service. During upper stage burnout, prior to This mechanical isolation system has demon- staging, the transient loads are relatively be- strated the capability to significantly alleviate nign. There are significant transient loads that the transient dynamic loads that occur during occur at both Stage 2 and Stage 3 ignition. flight. The isolation system can provide relief During the transient portion of these ignition to both the overall payload center of gravity events, the steady state axial loads are rela- loads and component or subsystem responses. tively nonexistent. Typically the system will reduce transient loads to approximately 50% of the level they would As dynamic response is largely governed be without the system. In addition, the system by payload characteristics, a mission specific generally reduces shock and vibration levels Coupled Loads Analysis (CLA) will be per- transmitted between the vehicle and space- formed, with customer provided finite element craft. The exact results can be expected to vary models of the payload, in order to provide more for each particular spacecraft and with loca- precise load predictions. Results will be refer- tion on the spacecraft. The isolation system enced in the mission specific ICD. For pre-

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200 400 600 800 1,000 1,200 lbm Payload Mass Does Not Include Random Vibe TM14025_051 Figure 4-3. Minotaur 3-Sigma High Maximum Acceleration as a Function of Payload Weight

Release 1.0 March 2002 4-2 Minotaur Payload User's Guide Section 4.0 - Payload Environment

does impact overall vehicle performance (by 10K approximately 20-40 lb [9-18 kg]) and the (10,000, available payload dynamic envelope (by up to 5K (1000, 3500) 3500) 4.0 in (10.16 cm) axially and approximately 1.0 2K (1850, (10,000, in (2.54 cm) laterally). 3000) 3000) 1K

4.2. Payload Vibration Environment 500 The in-flight random vibration curve shown in Figure 4-4 encompasses all flight vi- 200 bration environments. SRS (G's) 100 (100, 55)

1e-1 50 (100, 51) Separating Shock Breakpoints Non-Separating Shock Frequency PSD 20 (Hz) (g2/Hz) 10 20 0.002 100 200 300 500 1K 2K 3K 5K 10K 60 0.004 300 0.004 Frequency (Hz) 800 0.012 TM14025_052

/Hz) 1000 0.012 2 1e-2 2000 0.002 Figure 4-5. Maximum Shock Environment - 3.5344 gRMS 60 Sec Duration Launch Vehicle to Payload PSD (g

1e+4

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Frequency (Hz) TM14025_069 1e+3

/Hz) Breakpoints

Figure 4-4. Payload Random Vibration 2 Environment During Flight Natural Frequency SRS (G) 4.3. Payload Shock Environment PSD (g 1e+2 (Hz) Q=10 The maximum shock response spectrum 100 85 1000 3500 at the base of the payload from all launch ve- 1850 5000 hicle events will not exceed the flight limit 10,000 5000 levels in Figure 4-5 (separating shock). For 1e+1 missions that do not utilize an Orbital supplied 100 1000 10000 Frequency (Hz) payload separation system, the shock re- TM14025_070 sponse spectrum at the base of the payload Figure 4-6. Maximum Shock Enviroment - from vehicle events will not exceed the lev- Payload to Launch Vehicle els in Figure 4-5 (non-separating shock). 4.4. Payload Acoustic Environment If the payload employs a non-Orbital The acoustic levels during lift-off and separation system, then the shock delivered to the Stage 4 vehicle interface must not ex- powered flight will not exceed the flight limit ceed the limit level characterized in Figure levels shown in Figure 4-7. If the vehicle is 4-6. Shock above this level could require launched over a flame duct, the acoustic lev- requalification of components or an accep- els can be expected to be lower than shown. tance of risk by the payload customer. This has been demonstrated with flight data.

Release 1.0 March 2002 4-3 Minotaur Payload User's Guide Section 4.0 - Payload Environment

135 a. Design Limit Load — The maximum predicted ground-based, powered flight or

130 on-orbit load, including all uncertainties. b. Design Yield Load — The Design Limit

125 Load multiplied by the recommended Yield Factor of Safety (YFS) indicated in Figure 4-8. The payload structure must 120 have sufficient strength to withstand simul- taneously the yield loads, applied tem- 115 perature, and other accompanying environmental phenomena for each design 110 condition without experiencing detrimen- Sound Pressure Level (dB - re:2.9e-9psi) Sound Pressure Level tal yielding or permanent deformation. 105 10 100 1000 10000 c. Design Ultimate Load — The Design Limit Frequency (Hz) Load multiplied by the recommended Ul- Breakpoints Breakpoints Cont'd timate Factor of Safety (UFS) indicated in Natural Natural Figure 4-8. The payload structure must Frequency SPL (dB- Frequency SPL (dB- (Hz) re:2.9e-9psi) (Hz) re:2.9e-9psi) have sufficient strength to withstand simul- 20 113 500 128.6 taneously the ultimate loads, applied tem- 25 118 630 128 perature, and other accompanying envi- 32 123 800 126.3 40 123.8 1000 124.7 ronmental phenomena without experienc- 50 124.7 1250 123 ing any fracture or other failure mode of 63 125.5 1600 121.3 the structure. 80 126.3 2000 119.7 100 127.2 2500 118 125 128 3150 116.3 Design Factor of Safety on Limit Loads 160 128.8 4000 114.7 Design and Test 200 129.7 5000 113 Test Options Yield Ultimate Level 250 130.5 6300 111.3 (YFS) (UFS) 315 129.9 8000 109.7 400 129.3 10,000 108 Dedicated Test Article 1.00 1.25 UFS

TM14025_071 Proto-Flight Article 1.25 1.50 1.25

Figure 4-7. Payload Acoustic Environment Proof Test Each 1.10 1.25 1.10 During Liftoff and Flight Flight Article No Static Test 1.60 2.00 N/A

4.5. Payload Structural Integrity and Envi- TM14025_047 ronments Verification Figure 4-8. Factors of Safety Payload Design The primary support structure for the and Test spacecraft must possess sufficient strength, ri- gidity, and other characteristics required to 4.5.1. Recommended Payload Testing and survive the critical loading conditions that ex- Analysis ist within the envelope of handling and mis- Sufficient payload testing and/or analysis sion requirements, including worst case pre- must be performed to ensure the safety of dicted ground, flight, and orbital loads. It must ground crews and to ensure mission success. survive those conditions in a manner that as- The payload design should comply with the sures safety and that does not reduce the mis- testing and design factors of safety in Figure sion success probability. Spacecraft design 4-8. At a minimum, it is recommended that loads are defined as follows: the following tests be performed:

Release 1.0 March 2002 4-4 Minotaur Payload User's Guide Section 4.0 - Payload Environment

a. Structural Integrity — Static loads, sine The payload organization must provide vibration, or other tests should be per- Orbital with a list of the tests and test levels to formed that combine to encompass the which the payload was subjected prior to pay- acceleration load environment pre- load arrival at the integration facility. sented in Section 4.1. b. Random Vibration — The flight level 4.6. Thermal and Humidity Environments environment is defined in Figure 4-4. The thermal and humidity environment Recommended test levels are defined to which the payload may be exposed during in Figure 4-4. vehicle processing and pad operations are de- c. Acoustics — Depending on the pay- fined in the sections that follow and listed in load configuration, the payload or- Figure 4-10. ganization may elect to perform acoustic testing on the payload, or 4.6.1. Ground Operations sub-sections of the payload, in addi- The payload environment will be main- tion to, or in-lieu of, random vibra- tained by the Ground or Pad Air Conditioning tion testing. The acoustic levels are Systems (GACS or PACS). The GACS provides defined in Figure 4-7. conditioned air to the payload in the Vehicle Assembly Building (VAB) after fairing integra- Test tion. The PACS is used at the launch pad after Test TypePurpose Test Level vehicle stacking operations. Air Conditioning Random Vibration: the Qualification Flight Limit Level + 6dB (AC) is not provided during transport or lifting Flight Limit Level Is Acceptance Flight Limit Level operations without the enhanced option that Characterized in Figure 4-4 Protoflight Flight Limit Level + 3dB includes High Efficiency Particulate Air (HEPA) filtration. The conditioned air enters the fairing Figure 4-9. Recommended Payload Testing at a location forward of the payload, exits aft of Requirements the payload and is provided up to 5 minutes prior

Temp Range Purity Class Event Control Humidity (%) Deg C Deg F (Note 3) Pre-Payload Fairing Installation • Outside VAB Clean Tent 23 ±5 74 ±10 AC 45 ±15 None • Inside VAB Clean Tent 23 ±5 74 ±10 Filtered AC 45 ±15 100 K (M6.5) (Non Standard)

Post-Payload Fairing Installation PLF Inlet PLF Inlet (GSE) • VAB (GACS) 23 ±5 74 ±10 Filtered AC 45 ±15 100 K (M6.5) • Transportation to Launch Pad Ambient Ambient Filtered Ambient <60 100 K (M6.5) (Optional) (Note 4) (Note 4) (Note 1) • Lifting Operations (Optional) Ambient Ambient Filtered AC Ambient Optional

PACS (Ground) PLF Inlet PLF Inlet Filtered AC (Note 2) 100 K (M6.5) 13 - 29 55 - 85 Filtered AC Notes: 1. GSE AC Performance is Dependent Upon Ambient Conditions. Temperature Is Selectable and Controlled to Within ±4 ˚F (±2 ˚C) of Set Point. Resultant Relative Humidity is Maintained to 45 ±15%. 2. PACS Performance is Dependent Upon Ambient Conditions (Dew Point). Temperature is Selectable and Controlled Within ±4 ˚F (±2 ˚C) of Set Point. Resultant Relative Humidity is Maintained to 45 ±15%. 3. Class 10K (M5.5) Can Be Provided Inside the VAB Clean Tent and at the Payload Fairing Air-Conditioning Inlet on a Mission Specific Basis As a Non-Standard Service. 4. Temperature Control During Transport and Lifting Ops Available as a Non-Standard Service.

Figure 4-10. Payload Thermal and Humidity Environments

Release 1.0 March 2002 4-5 Minotaur Payload User's Guide Section 4.0 - Payload Environment to launch. Baffles are provided at the air condi- of the payload. As a non-standard service, a low tioning inlet to reduce impingement velocities emissivity coating can be applied to reduce on the payload if required. emissivity to less than 0.1. Interior surfaces aft of the payload interface will be maintained at Fairing inlet conditions are selected by the less than 250 °F (121 °C). Figure 4-11 shows customer, and are bounded as follows: the worst case transient temperature profile of • Dry Bulb Temperature: 55-85 °F the inner fairing surface adjacent to the pay- (13-29 °C) controllable to ±4 °F (±2 °C) load during powered flight. of setpoint; • Dew Point Temperature: 38-62 °F This temperature limit envelopes the (3-17 °C) maximum temperature of any component inside • Relative Humidity: determined by drybulb the payload fairing with a view factor to the and dewpoint temperature selections and payload with the exception of the Stage 4 mo- generally controlled to within ±15%. tor. The maximum Stage 4 motor surface tem- Relative humidity is bound by the psy- perature exposed to the payload will not ex- chrometric chart and will be controlled ceed 350 °F (177 °C), assuming no shielding be- such that the dew point within the fairing tween the aft end of the payload and the for- is never reached. ward dome of the motor assembly. Whether this temperature is attained prior to payload separa- 4.6.2. Powered Flight tion is dependent upon mission timeline. The maximum fairing inside wall temperature will be maintained at less than 200 °F The fairing peak vent rate is typically less (93 °C), with an emissivity of 0.92 in the region than 0.6 psi/sec. Fairing deployment will be

180 350

140 300 F) C) ˚ ˚ 250 100 200

60 150 Temperature ( Temperature ( 100 20 50

0 -20 0 25 50 75 100 125 150 Flight Time (Sec) • Data Analytically Derived (Using Flight-Verified Models) • Worst Case Heating Profile (Hot Trajectory)

• Fairing Inner Surface Temperature in Payload Region TM14025_053a

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

Release 1.0 March 2002 4-6 Minotaur Payload User's Guide Section 4.0 - Payload Environment initiated at a time in flight that the maximum load processing activities up to fairing encap- dynamic pressure is less than 0.01 psf. sulation. The soft walled clean room and anteroom(s) utilize HEPA filter units to filter the 4.6.3. Nitrogen Purge (non-standard service) air and hydrocarbon content is maintained at 15 If required for spot cooling of a payload ppm or less. The payload organization is respon- component, Orbital will provide GN2 flow to sible for providing the necessary clean room gar- localized regions in the fairing as a non-standard ments for payload staff as well as vehicle staff service. The GN2 will meet Grade B specifica- that need to work inside the clean room. tions, as defined in MIL-P-27401C and can be regulated to at least 5 scfm. The GN2 is on/off The inner surface of the entire surface of controllable in the launch equipment vault and the fairing and payload cone assemblies can in the launch control room. be cleaned to cleanliness criteria which ensures no particulate matter visible with normal vision The system’s regulators are set to a desired when inspected from 6 to 18 inches under 100 flow rate during prelaunch processing. The sys- ft-candle incident light. The same will be true tem cannot be adjusted after the launch pad has when the surface is illuminated using black light, been cleared of personnel. 3200 to 3800 Angstroms (Visibly Clean Plus Ul- traviolet). In addition, Orbital can ensure that Payload purge requirements must be co- all materials used within the encapsulated vol- ordinated with Orbital via the ICD to ensure that ume have outgassing characteristics of less than the requirements can be achieved. 1.0% TML and less than 0.1% CVCM. Items that don’t meet these levels can be masked to 4.7. Payload Contamination Control ensure they are encapsulated and will have no The Minotaur vehicle, all payload integra- significant effect on the payload. tion procedures, and Orbital’s contamination control program have been designed to minimize With the enhanced contamination control the payload’s exposure to contamination from option, Orbital provides an Environmental Con- the time the payload arrives at the payload pro- trol System (ECS) from payload encapsulation cessing facility through orbit insertion and sepa- through vehicle lift-off. The ECS continuously ration. The Vehicle Assembly Building is main- purges the fairing volume with clean filtered air. tained as a visibly clean, temperature and hu- Orbital’s ECS incorporates a HEPA filter unit to midity controlled work area at all times. The provide FED-STD-209 Class M5.5 (10,000) air. Minotaur assemblies that affect cleanliness within Orbital monitors the supply air for particulate the encapsulated payload volume include the matter via a probe installed upstream of the fair- fairing, and the payload cone assembly. These ing inlet duct prior to connecting the air source assemblies are cleaned such that there is no par- to the payload fairing. ticulate or non-particulate matter visible to the normal unaided eye when inspected from 2 to 4 Minotaur contamination control is based feet under 50 ft-candle incident light (Visibly on industry standard contamination reference Clean Level II). If required, the payload can be documents, including the following: provided with enhanced contamination control MIL-STD-1246C, “Product Cleanliness Levels as a non-standard service. and Contamination Control Program” FED-STD-209E, “Airborne Particulate Cleanliness With enhanced contamination control, a Classes in Cleanrooms and Clean Zones.” soft walled clean room can be provided to ensure a FED-STD-209 Class M6.5 (100,000) or 4.8. Payload Electromagnetic Environment Class M5.5 (10,000) environment during all pay- The payload Electromagnetic Environment

Release 1.0 March 2002 4-7 Minotaur Payload User's Guide Section 4.0 - Payload Environment

(EME) results from two categories of emitters: 1) cific EME experienced by the payload during Minotaur onboard antennas and 2) Range ra- ground processing at the VAB and the launch dar. All power, control and signal lines inside site will depend somewhat on the specific the payload fairing are shielded and properly facilities that are utilized as well as opera- terminated to minimize the potential for Elec- tional details. However, typically the field tromagnetic Interference (EMI). The Minotaur strengths experienced by the payload during payload fairing is Radio Frequency (RF) ground processing with the fairing in place opaque, which shields the payload from ex- are controlled procedurally and will be less ternal RF signals while the payload is encap- than 2 V/m from continuous sources and less sulated. Based on analysis and supported by than 10 V/m from pulse sources. The highest test, the fairing provides 20 db attenuation be- EME during powered flight is created by the tween 1 and 10000 MHz. C-Band transponder transmission which results in peak levels at the payload interface Figure 4-12 lists the frequencies and plane of 28 V/m at 5765 MHz. Range trans- maximum radiated signal levels from vehicle mitters are typically controlled to provide a antennas that are located near the payload field strength of 10 V/m or less. This EME during ground operations and powered flight. should be compared to the payload’s RF sus- Antennas located inside the fairing are inac- ceptibility levels (MIL-STD-461, RS03) to de- tive until after fairing deployment. The spe- fine margin.

SOURCE 12345 Function Command Tracking Tracking Launch Instrumentation Destruct Transponder Transponder Vehicle on Telemetry (Optional) Receive/Xmit Receive Transmit Receive Transmit Transmit Band UHF C-Band C-Band S-Band S-Band Frequency 416.5 or 5765 5690 2288.5 2269.5 (MHz) 425.0 Bandwidth N/A N/A Power Output N/A 400 W (peak) N/A 10 W 10 W Sensitivity -107 dB -70 dB Modulation Tone Pulse Code Pulse Code PCM/FM PCM/FM

TM14025_072 Figure 4-12. Minotaur Launch Vehicle RF Emitters and Receivers

Release 1.0 March 2002 4-8 Minotaur Payload User's Guide Section 5.0 - Payload Interfaces

5. PAYLOAD INTERFACES evaluate payload dynamic deflection with the This section describes the available me- Coupled Loads Analysis (CLA). The payload con- chanical, electrical and Launch Support Equip- tractor should assume that the interface plane is ment (LSE) interfaces between the Minotaur rigid; Orbital has accounted for deflections of the launch vehicle and the payload. interface plane. The CLA will provide final verification that the payload does not violate the dy- 5.1. Payload Fairing namic envelope. The standard payload fairing consists of two graphite composite halves, with a nosecap 5.1.2. Payload Access Door bonded to one of the halves, and a separation Orbital provides one 8.5 in. x 13.0 in. system. Each composite half is composed of a (21.6 cm x 33.0 cm), graphite, RF-opaque pay- cylinder and an ogive section. The two halves load fairing access door. The door can be posi- are held together by two titanium straps, both of tioned according to user requirements within the which wrap around the cylinder section, one near zone defined in Figure 5-2. The position of the its midpoint and one just aft of the ogive section. payload fairing access door must be defined no Additionally, an internal retention bolt secures later than L-8 months. Additional access doors the two fairing halves together at the surface can be provided as a non-standard service. where the nosecap overlaps the top surface of the other fairing half. The base of the fairing is 5.1.3. Increased Volume Payload Fairing separated using a frangible joint. Severing the An increased volume payload fairing is frangible joint allows each half of the fairing to currently in development for the Minotaur ve- then rotate on hinges mounted on the Stage 3 hicle. This larger fairing is discussed in more side of the interface. A contained hot gas gen- detail in Section 8.1.4. eration system is used to drive pistons that force the fairing halves open. All fairing deployment 5.2. Payload Mechanical Interface and Sepa- systems are non-contaminating. ration System Minotaur provides for a standard non- 5.1.1. Payload Dynamic Design Envelope The fairing drawing in Figures 5-1 show separating payload interface and several optional the maximum dynamic envelopes available for Orbital-provided payload separation systems. the payload during powered flight. The dynamic Orbital will provide all flight hardware and inte- envelopes shown account for fairing and vehicle gration services necessary to attach non-separat- structural deflections only. The payload contrac- ing and separating payloads to Minotaur. Ground tor must take into account deflections due to handling equipment is typically the responsibil- spacecraft design and manufacturing tolerance ity of the payload contractor. All attachment stack-up within the dynamic envelope. Proposed hardware, whether Orbital or customer provided, payload dynamic envelope violations must be must contain locking features consisting of lock- approved by OSP via the ICD. ing nuts, inserts or fasteners.

No part of the payload may extend aft of 5.2.1. Standard Non-Separating Mechanical In- the payload interface plane without specific OSP terface approval. These areas are considered stayout Figure 5-3 illustrates the standard, non- zones for the payload and are shown in Figure separating payload mechanical interface. This 5-1. Incursions to these zones may be approved is for payloads that provide their own separation on a case-by-case basis after additional verifica- system and payloads that will not separate. Di- tion that the incursions do not cause any detri- rect attachment of the payload is made on the mental effects. Vertices for payload deflection Avionics Structure with sixty #10 fasteners as must be given with the Finite Element Model to shown in Figure 5-3.

Release 1.0 March 2002 5-1 Minotaur Payload User's Guide Section 5.0 - Payload Interfaces

0° Pyrotechnic Event Harness Connector Pigtails Legend: to Payload Stayout Zone Clamp/Separation Payload System Components Stayout Zones

φ 77.7 30.6 Forward View ° ° 90 270 Looking Aft

Payload Interface Connector 180° 78.7 Payload φ ( 38 Inches 97 cm) Payload 31.0 Dynamic Separation System Envelope Stayout Zone

Notes: (1) Fairing Door Location Is Flexible Within a Specific Fairing Region. (Figure 5-2). R 270.5 (2) Payload Must Request Any 106.5 Envelope Aft of Payload Interface Plane. Ogive Mate (3) If Payload Falls within ACS Line Controllability Dead Band They Must Honor ACS 212.9 Stayout Zone. 83.8 (4) If the Payload Requires Nitrogen Cooling, then the Payload Envelope Will be Locally Reduced by 1 Inch Along the Cooling Tube Routing. 110.0 43.3 Payload Interface Plane φ 119.4 for Payload Separation 10.0 47.0 ACS Stayout System 4.0 Zone

Payload Interface Plane for Non-Separating Payloads 2.5 1.0

38 in (97 cm) Avionics Structure φ100.3 22 in (56 cm) Long 39.5 +X +Z Dimensions in cm in Side View TM14025_036

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

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Notes: 38" Payload 23" Payload Interface Place Interface Place Minotaur Coordinates Minotaur Station X* Minotaur Station X* +X (cm/in) (cm/in) Separable 1661.4/654.1 1685.5/663.6 Non-Separable 1651.5/650.2 1677.9/660.6 (1) Entire Access Hole Must Be Within Specified Range. +Z (2) One 8.5 in x 13.0 in (21.6 cm x 33.0 cm) Door per Mission Is Standard. (3) Edge of Door Cannot Be Within 5 in (13 cm) of Fairing Joints.

*Without a Soft Ride Load Isolation System

Station X +1,749.3 +688.7

Access Door Zone

Station X +1,669.8 +657.4

13 Arc 5 Length

13 Arc Length 5

Applies on Either Side of Fairing Joints at 0° and 180°

cm Dimensions in in TM14025_037

Figure 5-2. Payload Fairing Access Door Placement Zone

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22.9 9.0 Payload Harness 10.3 4.1 Minotaur Stage 4 Harness

45° Forward +X

Harness Access Hole ± φ 5.7 0.09 2.3 ±0.04 Forward Interface of φ 38 in (97 cm), VIEW A-A 56 cm (22 in) Long Avionics Structure Bolt Circle Consists of Rotated 90° CCW Applies at 45° (Pyrotechnic Event) and 0.20 in (60 0.51 cm) 225° (Payload Interface) Holes Equally Spaced, 98.6 Starting at 0° φ 0° 38.8 A Bolt Circle

A 45°

Pyrotechnic Event Connector

21.25 19.95 2X 2X Minotaur Coordinates 90° 270° +Y 1.90

+Z Payload Stayout Zone

119.4 φ 47.0 Fairing Dynamic Envelope 225° Payload Interface Connector

180° Forward View cm Looking Aft Dimensions in in TM14025_038 Figure 5-3. Non-Separable Payload Mechanical Interface

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5.2.2. Separating Mechanical Interface trical interface supports battery charging, ex- Three flight qualified optional separation ternal power, discrete commands, discrete te- systems are available, depending on payload lemetry, analog telemetry, serial communica- interface and size. The 38 in (97 cm) separable tion, payload separation indications, and up to payload interface is shown in Figure 5-4; the four redundant ordnance events. If an optional 23 in (59 cm) separable payload interface is Orbital-provided separation system is utilized, shown in Figure 5-5; the 17 in (43 cm) sepa- Orbital will provide all the wiring through the rable payload interface is shown in Figure 5-6. separable interface plane, as illustrated in Fig- Each of these three systems are based on a ure 5-8 and Figure B-1 of Appendix B. If the Marmon band design. option is not exercised the customer will be responsible to provide the wiring from the space- The separation ring to which the payload craft to the separation plane. attaches is supplied with through holes and the separation system is mated to the spacecraft during processing at the VAB. The weight of hard- 5.3.1. Payload Umbilical Interfaces ware separated with the payload is approxi- Orbital can provide a maximum of 60 mately 8.7 lbm (4.0 kg) for the 38 in (97 cm) wires from the ground to the spacecraft via a system, 6.0 lbm (2.7 kg) for the 23 in (59 cm) dedicated payload umbilical within the vehicle. system, and 4.7 lbm (2.1 kg) for the 17 in (43 This internal umbilical is approximately 25 ft in cm) system. Orbital-provided attachment bolts length. This umbilical is a dedicated pass through to this interface can be inserted from either the launch vehicle or the payload side of the inter- harness, which allows the payload command, face (NAS630xU, dash number based on pay- control, monitor, and power to be easily con- load flange thickness). The weight of the bolts, figured for user requirements. The closest prox- nuts, and washers connecting the separation imity for locating customer supplied payload system to the payload is allocated to the sepa- GSE equipment is the LEV (Launch Equipment ration system and included in the launch ve- Vault). The cabling from the LEV to the launch hicle mass. vehicle is approximately 350 ft.

At the time of separation, the payload is 5.3.2. Payload Battery Charging ejected by matched push-off springs with suffi- Orbital provides the capability for remote cient energy to produce the relative separation controlled charging of payload batteries, using velocities shown in Figure 5-7. If non-standard a customer provided battery charger. This separation velocities are needed, different power is routed through the payload umbilical springs may be substituted on a mission-specific cable. Up to 4.0 Amps per wire pair can be basis as a non-standard service. accommodated. The payload battery charger should be sized to withstand the line loss from 5.2.3. Orbital Supplied Drill Templates the LEV to the spacecraft. Orbital will provide a matched drill template to the payload contractor to allow accu- rate machining of the fastener holes. The Or- 5.3.3. Payload Command and Control bital provided drill template is the only approved Discrete sequencing commands gener- fixture for drilling the payload interface. The ated by the launch vehicle’s flight computer are payload contractor will need to send a contracts available to the payload as closed circuit opto- letter requesting use once they have determined isolator pulses in lengths of 40 ms multiples. The their need dates. current at the payload interface must be less than 10 mA. The payload must supply the volt- 5.3. Payload Electrical Interfaces age source and current limiting resistance to The existing design for the payload elec- complete this circuit.

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Minotaur Coordinates Bolt Cutters (2) (Redundant) +Y 0û +Z Payload Interface Payload Pyro Connector

90û 270û

Bolt Circle Consists of 0.19 in (60 0.48 cm) Payload Holes Equally Spaced, Push-Off Starting at 0û Springs (4 Places)

Payload Umbilical Connector

180û Forward View Looking Aft

4.0 kg (8.7 lbm) Remains with Payload (Includes Harness) 5.0 10.0 2.0 4.0 98.58 φ Bolt Circle 38.81 Payload Interface Separation Plane Plane

Payload Separation Clamp Band

+X Avionics Structure +Y

Side View cm Dimensions in in TM14025_039

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

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Payload Interface Payload Pyro Connector

0° Payload Push-Off Springs (4 Places) Clamp Band

Minotaur Bolt Cutters (2) Coordinates (Redundant) +Y 90° 270° +Z

Adpater Cone Bolt Circle Consists of 0.25 in (32 0.64 cm) Holes Equally Spaced, ° 180° Starting at 0

Payload Umbilical Connector

Forward View Looking Aft

3.75 7.49 2.7 Kg (6.00 lbm) 1.48 2.95 Remains with Payload φ 59.06 Bolt Circle (Includes Harness) 23.25 Payload Attachment Separation Plane Plane

Bolt Cutters (2) (Redundant)

Payload Separation Clamp Band Adapter Cone

cm +X Side View Dimensions in in

+Y TM14025_040

Figure 5-5. 23 in (59 cm) Separable Payload Interface

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0°

Bolt Cutters (2) (Redundant) Clamp Band Bolt Circle Consists of 0.25 in (24 0.64 cm) Holes Equally Spaced, Payload Push-Off Starting at 0¼ Springs (12 Places)

90° 270°

Minotaur Coordinates

Adapter Cone

180° Forward View Looking Aft

4.7 lbm (2.1 Kg) 3.75 φ 43.2 Remains with Payload 1.48 17.0 7.49 Bolt Circle 2.95 Payload Attachment Plane Separation Plane

Payload Separation Bolt Cutters (2) Clamp Band (Redundant)

Adapter Cone

cm Dimensions in +Y Side View from 0û in TM14025_041

Figure 5-6. 17 in (43 cm) Separable Payload Interface

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1.75

38 in (97 cm) Interface 1.5 5.00 23 in (59 cm) Interface 17 in (43 cm) Interface 1.25 4.00

1.00 3.00

.75 Separation Velocity (m/sec) Velocity Separation 2.00 (ft/sec) Velocity Separation .5 0 100 200 300 400 500 600 kg

0 200 400 600 800 1,000 1,200 lbm Payload Weight TM14025_042

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

Launch Vehicle Payload Spacecraft Interface Plane Plug Shell

Receptacle Shell Separation Plane S Socket Contacts Harness Recommend Length P Pin Contacts Hard Mount Specified by Payload Mate #2 Performed at VAB PS

Plug with S P Pin Contacts

Receptacle with Mate #1 Performed Socket Contacts Can Be Supplied at Orbital During to Payload Note: Sep System and Pigtails Delivered Separation System to VAB as a Unit Assembly TM14025_043

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

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5.3.4. Pyrotechnic Initiation Signals centerline for the standard configuration (within Orbital provides the capability to directly the accuracy listed in Figure 5-9). Payloads initiate 16 separate pyrotechnic conductors whose c.g. extend beyond the 1.5 in. lateral through two dedicated Ordnance Driver Mod- offset limit will require Orbital to verify the spe- ules (ODM). The ODM provides a 10 A, 100 cific offsets that can be accommodated. ms, current limited pulse into a 1 ±0.1 W initiation device. Measurement Accuracy

± ± 5.3.5. Payload Telemetry Mass 1 lbm ( 0.5 kg) Standard Minotaur service provides a Principal Moments of Inertia ±5% number of dedicated payload discrete (bi-level) and analog telemetry monitors through dedi- Cross Products of Inertia ±0.5 sl - ft2(±0.7 kg - m2) cated channels in the vehicle encoder. For dis- Center of Gravity X, Y and Z Axes ±6.4 mm (±0.25 in) crete monitors, the payload customer must provide the 5 Vdc source and the return path. The TM14025_045 current at the payload interface must be less Figure 5-9. Payload Mass Properties than 10 mA. Separation breakwire monitors can Measurement Tolerance be specified if required. The number of analog channels available for payload telemetry moni- 5.4.2. Final Mass Properties Accuracy toring is dependent on the frequency of the data. The final mass properties statement must Payload telemetry requirements and signal specify payload weight to an accuracy of at characteristics will be specified in the Payload least 1 lbm (0.5 kg), the center of gravity to an ICD and should not change once the final te- accuracy of at least 0.25 in (6.4 mm) in each lemetry format is released at approximately L- axis, and the products of inertia to an accuracy 6 months. of at least 0.5 slug-ft2 (0.7 kg-m2). In addition, if the payload uses liquid propellant, the slosh fre- 5.3.6. Non Standard Electrical Interfaces quency must be provided to an accuracy of 0.2 Non-standard services such as serial Hz, along with a summary of the method used command and telemetry interfaces can be ne- to determine slosh frequency. gotiated between OSP and the payload contractor on a mission-by-mission basis. 5.4.3. Pre-Launch Electrical Constraints Prior to launch, all payload electrical in- 5.3.7. Electrical Launch Support Equipment terface circuits are constrained to ensure there is Orbital will provide space for a rack of no current flow greater than 10 mA across the customer supplied EGSE in the LCR, or either of payload electrical interface plane. The primary the on-pad equipment vaults. The equipment support structure of the spacecraft shall be elec- will interface with the launch vehicle/space- trically conductive to establish a single point elec- craft through either the dedicated payload um- trical ground. bilical interface or directly through the payload access door. The payload customer is responsible for providing cabling from the EGSE loca- 5.4.4. Payload EMI/EMC Constraints tion to the launch vehicle/spacecraft. The Minotaur avionics share the payload area inside the fairing such that radiated emis- 5.4. Payload Design Constraints sions compatibility is paramount. OSP places The following sections provide design no firm radiated emissions limits on the payload constraints to ensure payload compatibility with other than the prohibition against RF transmis- the Minotaur system. sions within the payload fairing. Prior to launch, Orbital requires review of the payload radiated 5.4.1. Payload Center of Mass Constraints emission levels (MIL-STD-461, RE02) to verify Along the Y and Z axes, the payload c.g. overall launch vehicle EMI safety margin (emis- must be within 1.5 in (3.8 cm) of the vehicle sion) in accordance with MIL-E-6051. Payload

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RF transmissions are not permitted after fairing for Minotaur payloads. These are compliance mate and prior to an ICD specified time after documents and must be strictly followed. It is separation of the payload. An EMI/EMC analy- the responsibility of the customer to ensure that sis may be required to ensure RF compatibility. the payload meets all OSP, Orbital, and range imposed safety standards. Payload RF transmission frequencies must be coordinated with Orbital and range officials Customers designing payloads that employ to ensure non-interference with Minotaur and hazardous subsystems are advised to contact OSP range transmissions. Additionally, the customer early in the design process to verify compliance must schedule all RF tests at the integration site with system safety standards. with Orbital in order to obtain proper range clear- ances and protection.

5.4.5. Payload Dynamic Frequencies 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 12 Hz. However, this value is effected significantly by other factors such as incorporation of a spacecraft isolation system and/or separation system. Therefore, the final determination of compatibility must be made on a mission-specific basis.

5.4.6. Payload Propellent Slosh A slosh model should be provided to Or- bital in either the pendulum or spring-mass format. Data on first sloshing mode are required and data on higher order modes are desirable. The slosh model should be provided with the payload finite element model submittals.

5.4.7. Payload-Supplied Separation Systems If the payload employs a non-Orbital separation system, then the shock delivered to the Stage 4 vehicle interface must not exceed the limit level characterized in Section 4.3 (Figure 4-6). As stated in that section, shock above this level could require a requalification of components or an acceptance of risk by the payload customer.

5.4.8. System Safety Constraints OSP considers the safety of personnel and equipment to be of paramount importance. EWR 127-1 outlines the safety design criteria

Release 1.0 March 2002 5-11 Minotaur Payload User's Guide Section 6.0 - Mission Integration

6. MISSION INTEGRATION ordinate all mission planning and contracting activities. SMC Det 12/RP is supported by TRW 6.1. Mission Management Approach and other associate contractors for technical and The Minotaur program is managed logistical support, particularly utilizing their ex- through the US Air Force, Space and Missile Sys- tensive expertise and background knowledge of tems Center, Rocket Systems Launch Program the Minuteman booster and subsystems. (SMC Det 12/RP). SMC Det 12/RP serves as the primary point of contact for the payload custom- 6.1.2. Orbital Mission Responsibilities ers for the Minotaur launch service. A typical As the launch vehicle provider, Orbital’s integrated OSP organizational structure is shown responsibilities fall into four primary areas: in Figure 6-1. Open communication between a. Launch Vehicle Program Management SMC Det 12/RP, Orbital, and the customer, em- b. Mission Management phasizing timely transfer of data and prudent c. Engineering decision-making, ensures efficient launch ve- d. Launch Site Operations hicle/payload integration operations. Orbital assigns a Mission Manager to man- 6.1.1. SMC Det 12/RP Mission Reponsibilities age the launch vehicle technical and program- The program office for all OSP missions is matic interfaces for a particular mission. The Or- the SMC Det 12/RP. They are the primary Point bital Mission Manager is the single POC for all of Contact (POC) for all contractual and techni- aspects of a specific mission. This person has cal coordination. SMC Det 12/RP contracts with overall program authority and responsibility to Orbital to provide the Launch Vehicle and sepa- ensure that payload requirements are met and rately with commercial Spaceports and/or Gov- that the appropriate launch vehicle services are ernment Launch Ranges for launch site facilities provided. The Orbital Mission Manager will and services. Once a mission is identified, SMC jointly chair the Mission Integration Working Det 12/RP will assign a Mission Manager to co- Groups (MIWGs) with the SMC Det 12/RP Mis-

SMC Det 12/RP Rocket Systems Launch Program (RSLP) Office

Orbital Sciences Corp SMC Det 9 TRW OSP Prime Contractor RSLP VAFB Support RSLP SETA Support

• Vehicle Development • Range Interface • Technical Support • Mission and Payload • Facility Control • Independent Analysis Integration • Launch Support • MM Expertise • Launch Operations

Ogden ALC Spaceports and MM Motor Processing Government Ranges

• MM Motor Processing • Launch Facilities • Logistics • Range Instrumentation • X-Rays • Flight Safety TM14025_073 Figure 6-1. OSP Management Structure

Release 1.0 March 2002 6-1 Minotaur Payload User's Guide Section 6.0 - Mission Integration sion Manager. The Mission Managers respon- d. Range interface, safety, and flight sibilities include detailed mission planning, pay- operations activities, document ex- load integration services, systems engineering, changes, meetings and reviews. mission-peculiar design and analyses coordination, payload interface definition, launch range Figure 6-2 details the typical Mission Cycle coordination, integrated scheduling, launch site for a specific launch and how this cycle folds processing, and flight operations. into the Orbital vehicle production schedule with typical payload activities and milestones. A typi- 6.2. Mission Planning and Development cal Mission Cycle is based on a 18 month inter- OSP will assist the customer with mission val between mission authorization and launch. planning and development associated with This interval reflects the OSP contractual sched- Minotaur launch vehicle systems. These services ule and has been shown to be an efficient sched- include interface design and configuration con- ule based on Orbital’s Taurus and Pegasus pro- trol, development of integration processes, gram experience. However, OSP is flexible to launch vehicle analyses, facilities planning, negotiate either accelerated cycles, which take launch campaign planning to include range services and special operations, and integrated advantage of the Minotaur/Pegasus multi-cus- schedules. tomer production sets, or extended cycles required by unusual payload requirements, such The procurement, analysis, integration and as extensive analysis or complex payload-launch test activities required to place a customer’s pay- vehicle integrated designs or tests or funding limi- load into orbit are typically conducted over a 20 tations. month long standard sequence of events called the Mission Cycle. This cycle normally begins 6.3. Mission Integration Process 18 months before launch, and extends to eight weeks after launch. 6.3.1 Integration Meetings Once contract authority to proceed is re- The core of the mission integration pro- ceived, the Mission Cycle is initiated. The con- cess consists of a series of Mission Integration tract option designates the payload, launch date, and Range Working Groups (MIWG and RWG, and basic mission parameters. In response, the respectively). The MIWG has responsibility for Minotaur Program Manager designates an Or- all physical interfaces between the payload and bital Mission Manager who ensures that the the launch vehicle. As such, the MIWG creates launch service is supplied efficiently, reliably, and implements the Payload-to-Minotaur ICD in and on-schedule. addition to all mission-unique analyses, hardware, software, and integrated procedures. The The typical Mission Cycle interweaves the RWG is responsible for the areas of launch site following activities: operations; range interfaces; safety review and a. Mission management, document ex- approval; and flight design, trajectory, and guid- changes, meetings, and formal re- ance. Documentation produced by the RWG views required to coordinate and includes all required range and safety submit- manage the launch service. tals. b. Mission analyses and payload integration, document exchanges, and Working Group membership consists of meetings. the Mission Manager and representatives from c. Design, review, procurement, testing Minotaur engineering and operations organiza- and integration of all mission-pecu- tions, as well as their counterparts from the cus- liar hardware and software. tomer organization. While the number of meet-

Release 1.0 March 2002 6-2 Minotaur Payload User's Guide Section 6.0 - Mission Integration M21 M20 TM14025_074 M19 LRR FINAL PRE-LAUNCH M18 FINAL ILC FINAL (MEASURED) FINAL M17 QUICK-LOOK M16 MRR OD M15 FINAL P/L TO VAB FINAL FINAL M14 FINAL OR INTEGRATION PROCEDURES M13 FINAL PRELIM UPDATE FINAL M12 MDR-2 FINAL PL PRD Annex M11 CDR PRELIM REV PRELIM PRELIM M10 M9 FINAL M8 TEST PLAN M7 MDR-1 PSP M6 PRELIM M5 PRELIM M4 PRELIM M3 PRELIM SOC PRD Mission Integration Working Groups/ Operations Groups PRELIM PRELIM PRELIM M2 ATP PDR PI M1 Figure 6-2. Typical Minotaur Mission Integration Schedule PROCURE & FAB COMPONENTS INTEGRATION & TEST FIELD OPERATIONS PAYLOAD INTEGRATION LAUNCH OPERATIONS ATP MISSION KICK-OFF LAUNCH SITE SELECTION MISSION DESIGN REVIEWS ILC MILESTONES (TYPICAL) P/L DRAWINGS P/L FINITE ELEMENT MODEL P/L MASS PROPERTIES INTEGRATION DOCUMENTATION P/L ICD VERIFICATION PAYLOAD ICD COUPLED LOADS ANALYSIS MASS PROPS REPORT TARGETING RPT QUICK LOOK POST-FLIGHT REPORT FINAL POST FLIGHT REPORT MSPSP (Payload Annex) P/L HAZARDOUS PROCEDURES RANGE SAFETY TRAJECTORY PI/SOC (AS REQ'D) PRD/PSP(AS REQ'D) PL PRD Annex OR/OD MISSION CONSTRAINTS/COUNTDOWN LAUNCH VEHICLE MANUFACTURE Task Name MISSION INTEGRATION PAYLOAD MILESTONES AND DOCUMENTS MISSION DESIGN AND ENGINEERING SAFETY PROCESS RANGE DOCUMENTATION (UDS)

Release 1.0 March 2002 6-3 Minotaur Payload User's Guide Section 6.0 - Mission Integration ings, both formal and informal, required to de- which include—but are not limited to—the vari- velop and implement the mission integration pro- ous Range support agencies and U.S. Govern- cess will vary with the complexity of the space- ment agencies such as the U.S. Department of craft, quarterly meetings are typical. Transportation and U.S. State Department. Cus- tomer-provided documents represent the formal communication of requirements, safety data, sys- 6.3.2. Mission Design Reviews (MDR) tem descriptions, and mission operations plan- Two mission-specific design reviews will ning. The major products and submittal times be held to determine the status and adequacy of associated with these organizations are divided the launch vehicle mission preparations. They into two areas—those products that are provided are designated MDR-1 and MDR-2 and are typi- by the customer, and those produced by Or- cally held 6 months and 13 months, respectively, bital. after authority to proceed. They are each analo- gous to Preliminary Design Reviews (PDRs) and Critical Design Reviews (CDRs), but focus pri- 6.4.1. Customer-Provided Documentation marily on mission-specific elements of the launch Documentation produced by the cus- vehicle effort. tomer is detailed in the following paragraphs.

6.3.3. Readiness Reviews 6.4.1.1. Payload Questionnaire During the integration process, reviews are The Payload Questionnaire is designed held to provide the coordination of mission par- to provide the initial definition of payload re- ticipants and management outside of the regular quirements, interface details, launch site facili- contact of the Working Groups. Due to the vari- ties, and preliminary safety data to OSP. The ability in complexity of different payloads and customer shall provide a response to the Pay- missions, the content and number of these re- load Questionnaire form (Appendix A), or pro- views can be tailored to customer requirements. vide the same information in a different format, As a baseline, Orbital will conduct two readi- in time to support the Mission Kickoff Meeting. ness reviews as described below. The customer’s responses to the payload questionnaire define the most current payload re- Mission Readiness Review — Conducted quirements and interfaces and are instrumental within one month of launch, the Mission Readi- in Orbital’s preparation of numerous documents ness Review (MRR) provides a pre-launch assess- including the ICD, Preliminary Mission Analy- ment of integrated launch vehicle/payload/facil- sis, and launch range documentation. Addi- ity readiness prior to committing significant re- tional pertinent information, as well as prelimi- sources to the launch campaign. nary payload drawings, should also be included with the response. Orbital understands that a Launch Readiness Review — The Launch definitive response to some questions may not Readiness Review (LRR) is conducted at L-1 day be feasible. These items are defined during the and serves as the final assessment of mission normal mission integration process. readiness prior to activation of range resources on the day of launch. 6.4.1.2. Payload Mass Properties Payload mass properties must be pro- 6.4. Documentation vided in a timely manner in order to support Integration of the payload requires de- efficient launch vehicle trajectory development tailed, complete, and timely preparation and sub- and dynamic analyses. Preliminary mass prop- mittal of interface documentation. As the launch erties should be submitted as part of the MRD service provider, SMC Det 12/RP is the primary at launch vehicle authority to proceed. Up- communication path with support agencies, dated mass properties shall be provided at pre-

Release 1.0 March 2002 6-4 Minotaur Payload User's Guide Section 6.0 - Mission Integration defined intervals identified during the initial launches from other Ranges, a Range-specific mission integration process. Typical timing of PRD will be created. This document describes these deliveries is included in Figure 6-2. requirements needed to generally support the Minotaur launch vehicle. For each launch, an annex is submitted to specify the range support 6.4.1.3. Payload Finite Element Model needed to meet the mission’s requirements. This A payload mathematical model is required annex includes all payload requirements as well for use in Orbital’s preliminary coupled loads as any additional Minotaur requirements that may analyses. Acceptable forms include either a arise to support a particular mission. The cus- Craig-Bampton model valid to 120 Hz or a tomer completes all appropriate PRD forms for NASTRAN finite element model. For the final submittal to Orbital. coupled loads analysis, a test verified mathematical model is desired. 6.4.1.6.1. Launch Operations Requirements (OR) Inputs 6.4.1.4. Payload Thermal Model for Inte- To obtain range support for the launch op- grated Thermal Analysis eration and associated rehearsals, an OR must An integrated thermal analysis can be per- be prepared. The customer must provide all pay- formed for any payload as a non-standard ser- load pre-launch and launch day requirements for vice. A payload thermal model will be required incorporation into the mission OR. from the payload organization for use in Orbital’s integrated thermal analysis if it is required. The analysis is conducted for three mission phases: 6.5. Safety a. Prelaunch ground operations; b. Ascent from lift-off until fairing jetti- 6.5.1. System Safety Requirements son; and In the initial phases of the mission inte- c. Fairing jettison through payload de- gration effort, regulations and instructions that ployment. apply to spacecraft design and processing are reviewed. Not all safety regulations will apply to Models must be provided in SINDA for- a particular mission integration activity. Tailor- mat. There is no limit on model size although ing the range requirements to the mission unique turn-around time may be increased for large mod- activities will be the first step in establishing the els. safety plan. OSP has three distinctly different mission approaches affecting the establishment 6.4.1.5. Payload Drawings of the safety requirements: Orbital prefers electronic versions of pay- a. Baseline mission: Payload integration load configuration drawings to be used in the and launch operations are conducted mission specific interface control drawing, if at VAFB, CA possible. Orbital will work with the customer to b. Campaign/VAFB Payload Integration define the content and desired format for the mission: Payload integration is con- drawings. ducted at VAFB and launch operations are conducted from a non-VAFB 6.4.1.6. Program Requirements Document launch location. (PRD) Mission Specific Annex Inputs c. Campaign/Non-VAFB Payload Inte- To obtain range support, a PRD must be gration mission: Payload integration prepared. A Minotaur PRD has been submitted and launch operations are conducted and approved by the Western Range. For at a site other than VAFB.

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For the baseline and VAFB Payload Inte- radioactive materials, propellants, pressurized gration missions, spacecraft prelaunch operations systems, toxic materials, cryogenics, and RF ra- are conducted at Orbital’s VAB, Building 1555, diation. Procedures relating to these systems as VAFB. For campaign style missions, the space- well as any procedures relating to lifting opera- craft prelaunch operations are performed at the tions or battery operations should be prepared desired launch site. for safety review submittal. OSP will provide this information to the appropriate safety offices for Before a spacecraft arrives at the process- approval. ing site, the payload organization must provide the cognizant range safety office with certification that the system has been designed and tested in accordance with applicable safety requirements (e.g. EWR 127-1 Range Safety Require- ments for baseline and VAFB Payload Integra- tion missions). Spacecraft that integrate and/or launch at a site different than the processing site must also comply with the specific launch site’s safety requirements. Orbital will provide the customer coordination and guidance regarding applicable safety requirements.

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

6.5.2. System Safety Documentation For each Minotaur mission, OSP acts as the interface between the mission and Range Safety. In order to fulfill this role, OSP requires safety information from the payloader. For launches from either the Eastern or Western Ranges, EWR 127-1 provides detailed range safety regulations. To obtain approval to use the launch site facilities, specified data must be prepared and submitted to the OSP Program Office. This information includes a description of each payload hazardous system and evidence of compliance with safety requirements for each system. Drawings, schematics, and assembly and handling procedures, including proof test data for all lifting equipment, as well as any other information that will aid in assessing the respec- tive systems should be included. Major categories of hazardous systems are ordnance devices,

Release 1.0 March 2002 6-6 Minotaur Payload User's Guide Section 7.0 - Ground and Launch Ops

7. GROUND AND LAUNCH OPERATIONS 7.2.1.2. Vehicle Integration and Test Activi- ties 7.1. Minotaur/Payload Integration Overview The major vehicle components and sub- The processing of the Minotaur upper assemblies that comprise the Minotaur Upper stack utilizes many of the same proven tech- Stack Assembly, including the Stage 3 and Stage niques developed for the Pegasus and Taurus 4 Orion motors, are delivered to Orbital’s VAB launch vehicles. This minimizes the handling located at VAFB, CA. There, the vehicle is hori- complexity for both vehicle and payload. Hori- zontally integrated prior to the arrival of the pay- zontal integration of the Minotaur vehicle upper load. Integration is performed at a convenient stages simplifies integration procedures, increases working height, which allows relatively easy ac- safety and provides excellent access for the inte- cess for component installation, inspection and gration team. In addition, simple mechanical and test. electrical interfaces reduce vehicle/payload integration times, increase system reliability and mini- The integration and test process ensures mize vehicle demands on payload availability. that all vehicle components and subsystems are thoroughly tested. Since the Minuteman motors 7.2. Ground And Launch Operations are not available at the VAB, a high fidelity simu- Ground and launch operations are con- lator consisting of actual Minuteman components ducted in three major phases: is used. a. Launch Vehicle Integration — As- sembly and test of the Minotaur vehicle 7.2.1.2.1. Flight Simulation Tests b. Payload Processing/Integration — Flight Simulation Tests use the actual flight Receipt and checkout of the satellite software and simulate a “fly to orbit” scenario payload, followed by integration with using simulated Inertial Navigation System (INS) Minotaur and verification of inter- data. The Flight Simulation is repeated after each faces major change in vehicle configuration (i.e., Flight c. Launch Operations — Includes trans- Simulation #2 after stage mate, Flight Simulation port of the upper stack to the launch #3 with payload electrically connected (if re- pad, final integration, checkout, arm- quired) and Flight Simulation #4 after the pay- ing and launch. load is mechanically integrated). After each test, a complete review of the data is undertaken prior 7.2.1. Launch Vehicle Integration to proceeding. The payload nominally partici- pates in Flight Simulation #3 and #4.

7.2.1.1. Planning and Documentation Minotaur integration and test activities are 7.2.2. Payload Processing/Integration controlled by a comprehensive set of Work Pack- Payloads normally undergo initial check- ages (WPs), that describe and document every out and processing at Air Force or commercial aspect of integrating and testing Minotaur and facilities at VAFB. The payload is then sent to its payload. Mission-specific work packages are the VAB for integration with the Minotaur upper created for mission-unique or payload-specific stack. After arrival at the VAB, the payload com- procedures. All procedures are issued by pletes its own independent verification and Orbital’s Vandenberg Planning Department be- checkout prior to beginning integrated process- fore work is started. Any discrepancies encoun- ing with Minotaur. Following completion of tered are recorded on a Discrepancy Report and Minotaur and payload testing the payload will dispositioned as required. All activities are in be enclosed inside the fairing. The required pay- accordance with Orbital’s ISO 9001 certification. load environments will be maintained inside the

Release 1.0 March 2002 7-1 Minotaur Payload User's Guide Section 7.0 - Ground and Launch Ops fairing until launch. Any payload specific haz- 7.2.2.4. Final Processing and Fairing Closeout ardous procedures should be coordinated After successful completion of Flight Simu- through Orbital to the launch range no later than lation #4, all consumables are topped off and 120 days prior to first use (draft) and 30 days prior ordnance is connected. Similar payload opera- to first use (final). tions may occur at this time. Once consumables are topped off, final vehicle / payload closeout is 7.2.2.1. Payload to Minotaur Integration performed and the fairing is installed. The pay- The integrated launch processing activi- load will coordinate with OSP access to the pay- ties are designed to simplify final launch pro- load from payload mate until final closeout be- cessing while providing a comprehensive veri- fore launch. fication of the payload interface. The systems integration and test sequence is engineered to 7.2.2.5. Payload Propellant Loading ensure all interfaces are verified. Payloads utilizing integral propulsion systems with propellants such as hydrazine can be 7.2.2.2. Pre-Mate Interface Testing loaded and secured through coordinated Orbital If required, the electrical interface be- and contractor arrangements for use of the pro- tween Minotaur and the payload is verified pellant loading facilities in the VAB. This is a using a mission unique Interface Verification non-standard service. Test (IVT) to jointly verify that the proper function of the electrical connections and com- 7.2.2.6. Final Vehicle Integration and Test mands. These tests, customized for each mis- Due to operational constraints, the Lower sion, typically check bonding, electrical com- Stack Assembly, consisting of the Minuteman mo- patibility, communications, discrete com- tors, is processed by the Air Force at a separate mands and any off nominal modes of the pay- facility. After testing by Orbital, it is delivered load. After completing the IVT, a Flight Simu- directly to the launch pad to await the arrival of lation (Flight Sim #3) is performed with the the upper stack. After the vehicle is fully stacked payload electrically - but not mechanically - at the pad, final tests are completed to verify ve- connected to Minotaur to demonstrate the full hicle integrity and all interfaces to the range are sequence of events in a simulated flight sce- exercised. nario. Once Flight Sim #3 is successfully completed, the payload is mechanically mated to 7.3 Launch Operations the launch vehicle. For payloads with simplified or no electrical interfaces to Minotaur, it 7.3.1. Launch Control Organization may be acceptable to proceed to payload mate The Launch Control Organization is split immediately after the IVT. For pre-mate veri- into two groups: the Management group and the fication of the mechanical interface, the sepa- Technical group. The Management group con- ration system can also be made available be- sists of senior range personnel and Mission Di- fore final payload preparations. rectors/Managers for the launch vehicle and payload. The Technical Group consists of the per- 7.2.2.3. Payload Mating and Verification sonnel responsible for the execution of the launch Following the completion of Flight Sim #3, operation and data review/assessment for the the jumpers between the payload and Minotaur Payload, the Launch Vehicle and the Range. The are removed. Once the payload aft end close- Payload’s members of the technical group are outs are completed, the payload will be both me- engineers who provide technical representation chanically and electrically mated to the Minotaur. in the control center. The Launch Vehicle’s Following mate, the flight vehicle is ready for the members of the technical group are engineers final integrated systems test, Flight Simulation #4. who prepare the Minotaur for flight, review and

Release 1.0 March 2002 7-2 Minotaur Payload User's Guide Section 7.0 - Ground and Launch Ops assess data that is displayed in the Launch Con- trol 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 main- tain and monitor the voice and data equipment , tracking facilities and all assets involved with RF communications with the launch vehicle. In addition, the Range provides personnel responsible for the Flight Termination System monitoring and commanding.

Release 1.0 March 2002 7-3 Section 8.0 - Optional Enhanced Minotaur Payload User's Guide Capabilities

8. OPTIONAL ENHANCED CAPABILITIES nent or subsystem responses. Typically the sys- The OSP launch service is structured to tem will reduce transient loads to approximately provide a baseline vehicle configuration which 50% of the level they would be without the sys- is then augmented with optional enhancements tem. The exact results can be expected to vary to meet the unique needs of individual payloads. for each particular spacecraft and with location The baseline vehicle capabilities are defined in on the spacecraft. Generally, a beneficial reduc- the previous sections and the optional enhanced tion in shock and vibration will also be provided. capabilities are defined below. The isolation system does impact overall vehicle performance (by approximately 20-40 lb [9-18 8.1. Mechanical Interface and Separation Sys- kg]) and the available payload dynamic enve- tem Enhancements lope by up to 4.0 in (10.16 cm) axially and up to 1.0 in (2.54 cm) laterally. 8.1.1. Separation Systems 8.1.4. Increased Payload Volume Orbital offers several alternate separating To accommodate payloads larger than interfaces as optional services. The baseline op- those that can be accommodated by the stan- tional separation system is the Orbital 38 in. design, but several flight-proven, Orbital-developed Payload Attach Fittings (PAFs) can be utilized. These systems are discussed in Section 5.2.1 and 5.2.2. In addition, other separation systems rang- ing from an Orbital-developed 10 in. diameter system to the SAAB 38.81 in. system can be provided as additional negotiated options.

8.1.2. Additional Fairing Access Doors OSP provides one 8.5 in x 13.0 in (21.6 cm x 33.0 cm), graphite, RF-opaque payload fairing access door (per Section 5.1.2). Additional doors can be provided within the range specified in Section 5.1.2. Other fairing access configurations, such as small circular access panels, can also be provided as negotiated mission-specific enhancements.

8.1.3. Payload Isolation System OSP offers a flight-proven payload isolation system as a non-standard service. This Softride for Small Satellites (SRSS) was developed by Air Force Research Laboratory (AFRL) and CSA Engineering. It was successfully demonstrated on the two initial Minotaur missions in the typical configuration shown in Figure 8-1. This mechanical isolation system has demonstrated the capability to significantly alleviate the transient dynamic loads that occur during flight. The isolation system can provide relief to both the over- Figure 8-1. Softride for Small Satellites (SRSS) all payload center of gravity loads and compo- Payload Isolation System

Release 1.0 March 2002 8-1 Section 8.0 - Optional Enhanced Minotaur Payload User's Guide Capabilities

by approximately 50 to 250 lbm, depending on the orbit. Specific performance capability as- 175.0 sociated with the HAPS can be provided by con- tacting the OSP program office. HAPS is more 162.3 effective at higher altitude and also permits injection of shared payloads into different orbits. 22.7 HAPS, which is mounted inside the Avionics Structure, consists of a hydrazine propulsion 133.5 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 perfor- 35.6 mance and precise orbit injection. The HAPS propulsion subsystem consists of a centrally AVAILABLE mounted tank containing approximately 130 ENVELOPE 82.2 lbm (59 kg) of hydrazine, helium pressurization (DYNAMIC) gas, and three fixed, axially pointed thrusters. The hydrazine tank contains an integral blad- der which will support multiple restarts.

8.3. Environmental Control Options 54.7

8.3.1. Conditioned Air Conditioned air can be provided within INTERFACE PLANE 0.0 the fairing volume using an Environmental Con- trol System (ECS) via a duct that is retracted in

TM14025_075 the last minutes prior to launch. Temperature and humidity can be regulated within the limits Figure 8-2. Optional 61 in. Diameter Fairing indicated in Figure 4-10. A filter is installed to provide a Class 100,000 environment, typically. dard Pegasus-based fairing, a larger fairing de- Upon exercise of the Enhanced Cleanliness op- sign is being developed in conjunction with tion, a certified HEPA filter is used in the input AFRL. A preliminary dynamic envelope is duct to assure the necessary low particulate en- shown in Figure 8-2. A slight performance de- vironment (Class 100,000 or Class 10,000). crease will be incurred. 8.3.2. Nitrogen Purge 8.2. Performance Enhancements OSP can provide gaseous nitrogen purges to the payload after fairing encapsulation until 8.2.1. Insertion Accuracy lift-off. The system distribution lines are routed Enhanced insertion accuracy or support along the inner surface of the fairing. If required for multiple payload insertion can be provided for spot cooling of a payload component, Or- as an enhanced option utilizing the Hydrazine bital will provide GN2 flow to localized regions Auxiliary Propulsion System (HAPS). Orbital in- in the fairing. The GN2 will meet Grade B speci- sertion accuracy can typically be improved to fications, as defined in MIL-P-27401C and can ±10 nm (±18.5 km) or better in each apse and be regulated between 2.4-11.8 l/sec (5-25 scfm).

±0.1 deg in inclination. For orbits above 600 The GN2 is on/off controllable in the launch km, the HAPS can also increase payload mass equipment vault and in the launch control room.

Release 1.0 March 2002 8-2 Section 8.0 - Optional Enhanced Minotaur Payload User's Guide Capabilities

The system’s regulators are set to a de- characteristics of less than 1.0% TML and less sired flow rate during prelaunch processing. than 0.1% CVCM. Items that don’t meet these The system cannot be adjusted after the launch levels can be masked to ensure they are encap- pad has been cleared of personnel. sulated and will have no significant effect on the payload. Payload purge requirements must be coordinated with Orbital via the ICD to ensure that 8.3.3.3 High Cleanliness Fairing Environment the requirements can be achieved. With the enhanced contamination control option, Orbital provides an ECS from payload encapsulation until just prior to vehicle lift-off. 8.3.3. Enhanced Contamination Control The ECS continuously purges the fairing volume Understanding that some payloads have with clean filtered air. Orbital’s ECS incorpo- requirements for enhanced cleanliness, OSP of- rates a HEPA filter unit to provide FED-STD-209 fers a contamination control option, which is Class M5.5 (10,000) air. Orbital monitors the composed of the elements in the following sec- supply air for particulate matter via a probe in- tions (which is also discussed in Section 4.7). stalled upstream of the fairing inlet duct prior to Minotaur customers can also coordinate combi- connecting the air source to the payload fairing. nations of the elements listed below to meet the unique needs of their payloads. 8.4. Enhanced Telemetry Options 8.3.3.1 High Cleanliness Integration Environ- OSP can provide mission specific instru- ment (Class 10K or 100K) mentation and telemetry components to support With enhanced contamination control, a additional payload or experiment data acquisi- soft walled clean room can be provided to en- tion requirements. Telemetry options include ad- sure a FED-STD-209 Class M6.5 (100,000) or ditional payload-dedicated bandwidth and GPS- Class M5.5 (10,000) environment during all pay- based precision navigation data. load processing activities up to fairing encapsulation. The soft walled clean room and 8.4.1. Enhanced Telemetry Bandwidth anteroom(s) utilize HEPA filter units to filter the A second telemetry data stream capable air and hydrocarbon content is maintained at 15 of up to 2 Mbps data rate can be provided. Maxi- ppm or less. The payload organization is respon- mum data rates depend on the mission cover- sible for providing the necessary clean room gar- age required and the launch range receiver ments for payload staff as well as vehicle staff characteristics and configuration. This capa- that need to work inside the clean room. bility was successfully demonstrated on the inaugural Minotaur mission. 8.3.3.2 Fairing Surface Cleanliness Options The inner surface of the fairing and payload cone assemblies can be cleaned to cleanli- 8.4.2. Enhanced Telemetry Instrumentation ness criteria which ensures no particulate matter To support the higher data rate capability visible with normal vision when inspected from in Section 8.4.1, enhanced telemetry instrumen- 6 to 18 inches under 100 ft-candle incident light. tation can be provided. The instrumentation can The same will be true when the surface is illumi- include strain gauges, temperature sensors, ac- nated using black light, 3200 to 3800 Angstroms celerometers, analog data, and digital data con- (Visibly Clean Plus Ultraviolet). In addition, Or- figured to mission-specific requirements. This bital can ensure that all materials used within capability was successfully demonstrated on the the encapsulated volume have outgassing inaugural Minotaur mission.

Release 1.0 March 2002 8-3 Section 8.0 - Optional Enhanced Minotaur Payload User's Guide Capabilities

8.4.3. Navigation Data Precision navigation data using an independent GPS-receiver and telemetry link is available as an enhanced option. This option utilizes the Orbital-developed GPS Position Beacon (GPB) and provides a better than 100 m position accuracy with a nominal 1 Hz data rate. This capability was successfully demonstrated on the inaugural Minotaur mission.

Release 1.0 March 2002 8-4 Section 9.0 - Shared Launch Minotaur Payload User's Guide Accommodations

9. SHARED LAUNCH ACCOMMODATIONS Minotaur is uniquely capable of providing launches of multiple satellite payloads, le- Fairing veraging SMC Det 12/RP and Orbital’s exten- Dynamic sive experience in integrating and launching Envelope multiple payloads. In addition to the five satellites successfully separated on the inaugural Typical Minotaur JAWSAT mission, multiple spacecraft Forward configurations have been flown on many of Spacecraft Orbital’s Pegasus and Taurus missions to date. Volume φ 23.00 in. (2.95) Separation Ring Two technical approaches are available for accommodating multiple payloads. These Typical Aft Load Bearing design approaches are: Spacecraft Volume Load-Bearing Spacecraft — aft spacecraft φ 38 in. Separation Ring designed to provide the structural load (3.95) path between the forward payload and the φ 38 in. Avionics launch vehicle, maximizing utilization of Thrust Tube (22.00 in. Long) available mass performance and payload φ 46.00 fairing volume (Dynamic) TM14025_077

Non Load-Bearing Spacecraft — aft Figure 9-1. Typical Load Bearing Spacecraft spacecraft whose design cannot provide Configuration the necessary structural load path for the forward payload are those involving mechanical and electrical compatibility with the forward payload. Struc- 9.1. Load-Bearing Spacecraft tural loads from the forward payload during all Providing a load-bearing aft payload maxi- flight events must be transmitted through the aft mizes use of available volume and mass. The payload to the Minotaur. Orbital will provide available mass for the aft payload is determined minimum structural interface design criteria for by the Minotaur performance capability to orbit shear, bending moment, axial and lateral loads, less the forward payload and attach hardware and stiffness. mass. All remaining mission performance, ex- cluding a stack margin, is available to the aft The second approach involves the use of payload. The load-bearing spacecraft interfaces an Orbital design using the MicroStar bus, suc- directly to the avionics assembly interface and cessfully developed and flown for ORBCOMM the forward payload via pre-determined inter- spacecraft. The MicroStar bus features a circular faces. These interfaces include standard Orbital design with an innovative, low-shock separation separation systems and pass-through electrical system. The spacecraft bus is designed to allow connectors to service the forward payload. Fig- stacking of co-manifested payloads in “slices” ure 9-1 illustrates this approach. within the fairing. The bus design is compact and provides exceptional lateral stiffness. Two approaches may be taken for load- bearing spacecraft. The first approach is to use a A variation on the load bearing spacecraft design developed by other spacecraft suppliers, approach was flown on the JAWSAT inaugural which must satisfy Minotaur and forward pay- mission, in which the primary payload (JAWSAT) load structural design criteria. The principal re- was a Multiple Payload Adapter (MPA) from quirements levied upon load-bearing spacecraft which four small satellites were separated (Fig-

Release 1.0 March 2002 9-1 Section 9.0 - Shared Launch Minotaur Payload User's Guide Accommodations

9.2. Non Load-Bearing Spacecraft For aft spacecraft that are not designed for withstanding and transmitting structural loads from the forward payload, several options are possible including AFRL and Orbital design con- cepts, as well as the flight-proven Dual Payload Attach Fitting (DPAF).

The DPAF structure (Figure 9-4) is an all graphite structure which provides independent load paths for each satellite. The worst-case “design payload” for the DPAF is a 425 lbm (193 TM14025_078 kg) spacecraft with a 20 in. (51 cm) center of mass offset and first lateral frequency of typically Figure 9-2. JAWSAT Multiple Payload 12 Hz. The DPAF is designed to accommodate Adapter Load Bearing Spacecraft this “design payload” at both the forward and aft locations, although the combined mass of the two ure 9-2). After separating the smaller “piggy- payloads cannot exceed Minotaur capabilities. back” satellites, the JAWSAT MPA was also The upper spacecraft loads are transmitted separated as an autonomous satellite by utiliz- around the lower spacecraft via the DPAF struc- ing the Orbital 23 in. separation system and ture, thus avoiding any structural interface be- adapter cone. An updated concept to provide tween the two payloads. greater payload options and primary payload volume (by mounting directly to the avionics assembly) is shown in Figure 9-3.

Integrated coupled loads analyses will be performed with test verified math models provided by the payload. These analyses are required to verify the fundamental frequency and deflections of the stack for compliance with the Minotaur requirement of 12 Hz minimum. De- sign criteria provided by OSP will include “stack” margins to minimize interactive effects associated with potential design changes of each payload. OSP will provide the necessary engineering coordination between the spacecraft and launch vehicle.

Electrical pass-through harnesses will also need to be provided by the aft payload along with provisions for connectors and interface verification. The spacecraft supplier will need to provide details of the appropriate analyses and tests to OSP to verify adequacy of margins and show TM14025_079 that there is no impact to the forward spacecraft Figure 9-3. Five Bay Multiple Payload Adapter or the launch vehicle. Concept

Release 1.0 March 2002 9-2 Section 9.0 - Shared Launch Minotaur Payload User's Guide Accommodations

cm Primary Payload Dimensions in in Separation Plane Secondary Payload Ogive Radius Separation Plane 58 Separation System 269.2 23 106.0 128.8 97 Separation 43 Separation 50.7 38 System 17 System

φ 114.3 Available Primary φ 66.0 45.0 φ 76.0 Secondary Payload 26.0 Dynamic 29.9 Payload Envelope Volume Volume

101.6 55.9 +Y 40.0 22.0 Minotaur Avionics 102.9 40.5 Beginning of Ogive Adapter Cone Separation Plane +X TM14025_082

Figure 9-4. Dual Payload Attach Fitting Configuration

Release 1.0 March 2002 9-3 Minotaur Payload User's Guide Appendix A

Appendix A Payload Questionnaire

Release 1.0 March 2002 A-1 MINOTAUR PAYLOAD QUESTIONNAIRE

SATELLITE IDENTIFICATION FULL NAME:

ACRONYM:

OWNER/OPERATOR:

INTEGRATOR(s):

ORBIT INSERTION REQUIREMENTS*

SPHEROID q Standard (WGS-84, Re = 6378.137 km) q Other: ALTITUDE Insertion Apse: Opposite Apse: q nmi q nmi ___ ± __ q km ± q km or... Semi-Major Axis: Eccentricity: q nmi ___ ± __ q 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):

1 MINOTAUR PAYLOAD QUESTIONNAIRE

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):

2 MINOTAUR PAYLOAD QUESTIONNAIRE

SPACECRAFT PHYSICAL DIMENSIONS STOWED Length/Height: Diameter: CONFIGURATION q in q cm q in q cm

Other Pertinent Dimension(s):

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

ON-ORBIT 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: q lbm-in q kg-m

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

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

3 MINOTAUR PAYLOAD QUESTIONNAIRE

ASCENT TRAJECTORY REQUIREMENTS

Free Molecular Heating at Fairing Separation: q Btu/ft2/hr (Standard Service: 360 Btu/ft2/hr) FMH q W/m2 Fairing Internal Wall Temperature q deg F (Standard Service: 200°F) T q deg C 2 Dynamic Pressure at Fairing Separation: q lbf /ft 2 2 (Standard Service: 0.01 lbf /ft ) q q N/m 2 Ambient Pressure at Fairing Separation: q lbf /in (Standard Service: 0.3 psia) P q N/m2 2 Maximum Pressure Decay During Ascent: q lbf /in /sec (Standard Service: 0.6 psia) ÃÃÃÃÃÃÃÃÃÃÃÃÃÃÃÃ Ã3ÃÃ q 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 q deg F q deg C Ground/Launch Operations: Min q deg F q deg C (Standard Service is 55°F to 80°F) Component(s) Driving Temperature Constraint: Approximate Location(s):

HUMIDITY Relative Humidity: or,Dew Point: Max % Max q deg F q deg C Min % Min q deg F q deg C (Standard Service is 37 deg F) 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

4 MINOTAUR PAYLOAD QUESTIONNAIRE

ELECTRICAL INTERFACE Bonding Requirements:

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

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

If Yes, describe:

Other Electrical Requirements:

Please complete attached sheet of required pass-through signals.

RF RADIATION Time After Separation Until RF Devices Are Activated:

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

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

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

5 MINOTAUR PAYLOAD QUESTIONNAIRE

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

6 MINOTAUR PAYLOAD QUESTIONNAIRE

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 q in q cm Clocking (deg), Describe:

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

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

Axial Hz Lateral Hz

Recommended: > 35 Hz > 12 Hz OTHER Other Mechanical Interface Requirements:

7 MINOTAUR PAYLOAD QUESTIONNAIRE

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 SEB. [Note: Space limitations exist in the SEB, 350 ft umbilical cable length to spacecraft typical] Equipment Name / Type Approximate Size (LxWxH) Power Requirements

......

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

......

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

8 Minotaur Payload User's Guide Appendix B

APPENDIX B Electrical Interface Connectors

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

Release 1.0 March 2002 B-1 Minotaur Payload User's Guide Appendix B

Launch Vehicle Spacecraft Standard Quantity

Discrete Commands 5

5V Discrete Telemetry Monitors (5 Hz) 5

Separation Indicator 3

10 Amp, 100 msec Pyro Initiation 4 Redundant Pairs

Analog Telemetry 20 mV to 10 V pp pp 6 ±50 V

Serial Communication Link 1

RS-422/RS-485

60 Conductors GSE Pass-Through Wiring (Various AWG/Shielding Available) Note: Channels Can be Reallocated Based on Customer Requirements TM14025_080

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

Release 1.0 March 2002 B-2