Minotaur Iv Launch Vehicle Users Guide
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ORBITAL SCIENCES CORPORATION
January 2006
Minotaur IV User’s Guide
Release 1.1
Approved for Public Release Distribution Unlimited
Copyright 2006 by Orbital Sciences Corporation. All Rights Reserved.
Minotaur IV User’s Guide
Revision Summary
REVISION SUMMARY VERSION DOCUMENT
DATE
CHANGE
PAGE
1.0
TM-17589
Jan 2005
Initial Release
All
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TM-17589A
Jan 2006
General nomenclature, history, and administrative updates (no technical updates)
All
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•
Updated launch history
•
Corrected contact information
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Minotaur IV User’s Guide
Preface
This User’s Guide is intended to familiarize payload mission planners with the capabilities of the Orbital Suborbital Program ll (OSP-2) Minotaur IV Space Launch Vehicle (SLV) launch service. The information provided in this user’s guide is for initial planning purposes only. Information for development/design is determined through mission specific engineering analyses. The results of these analyses are documented in a mission-specific Interface Control Document (ICD) for the payloader organization to use in their development/design process. This document provides an overview of the Minotaur IV system design and a description of the services provided to our customers.
USAF SMC Det 12/RP 3548 Aberdeen Ave SE Kirtland AFB, NM 87117-5778 Telephone: Fax:
(505) 846-8957 (505) 846-5152
Additional copies of this User's Guide and Technical information may also be requested from Orbital at: Orbital Suborbital Program - Mission Development Orbital Sciences Corporation Launch Systems Group 3380 S. Price Road Chandler, AZ 85248 Telephone: E-mail:
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Preface
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Table of Contents
PAGE 1.
INTRODUCTION................................................................................................................................. 1
2.
MINOTAUR IV LAUNCH SERVICE .................................................................................................... 3 2.1.
Minotaur IV Launch System Overview.......................................................................................... 3
2.2.
Minotaur IV Launch Service.......................................................................................................... 3
2.3.
Minotaur IV Launch Vehicle .......................................................................................................... 4
2.3.1.
Stage 1, 2 and 3 Booster Assemblies....................................................................................... 4
2.3.2.
Stage 4 Booster/Avionics Assembly ......................................................................................... 5
2.3.2.1.
Avionics............................................................................................................................... 5
2.3.2.2.
Attitude Control Systems .................................................................................................... 5
2.3.2.3.
Telemetry Subsystem ......................................................................................................... 7
2.3.2.4.
Payload Fairing ................................................................................................................... 7
2.4.
Launch Support Equipment .......................................................................................................... 8
2.4.1. 3.
Transportable LSE Shelters ...................................................................................................... 8
GENERAL PERFORMANCE ............................................................................................................ 11 3.1.
Mission Profiles........................................................................................................................... 11
3.2.
Launch Sites ............................................................................................................................... 11
3.2.1.
Western Launch Sites ............................................................................................................. 11
3.2.2.
Eastern Launch Sites .............................................................................................................. 11
3.2.3.
Alternate Launch Sites ............................................................................................................ 11
3.3.
Performance Capability............................................................................................................... 11
3.4.
Injection Accuracy....................................................................................................................... 23
3.5.
Payload Deployment................................................................................................................... 23
3.6.
Payload Separation..................................................................................................................... 23
3.7.
Collision/Contamination Avoidance Maneuver ........................................................................... 23
4.
PAYLOAD ENVIRONMENT.............................................................................................................. 25 4.1.
Steady State and Transient Acceleration Loads......................................................................... 25
4.1.1.
Transient Loads....................................................................................................................... 26
4.1.2.
Steady-State Acceleration....................................................................................................... 26
4.2.
Payload Vibration Environment .................................................................................................. 26
4.2.1.
Random Vibration ................................................................................................................... 28
4.2.2.
Sine Vibration .......................................................................................................................... 28
4.3.
Payload Acoustic Environment ................................................................................................... 28
4.4.
Payload Shock Environment....................................................................................................... 28
4.5.
Payload Structural Integrity and Environments Verification........................................................ 32
4.5.1. 4.6.
Recommended Payload Testing and Analysis ....................................................................... 33 Thermal and Humidity Environments.......................................................................................... 33
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4.6.1.
Ground Operations.................................................................................................................. 33
4.6.2.
Powered Flight ........................................................................................................................ 34
4.6.3.
Nitrogen Purge (Non-Standard Service) ................................................................................. 35
4.7.
Payload Contamination Control .................................................................................................. 35
4.8.
Payload Electromagnetic Environment....................................................................................... 35
5.
PAYLOAD INTERFACES.................................................................................................................. 37 5.1.
Payload Fairing ........................................................................................................................... 37
5.1.1.
Payload Dynamic Design Envelope ........................................................................................ 37
5.1.2.
Payload Access Door.............................................................................................................. 37
5.2.
Payload Mechanical Interface and Separation System .............................................................. 37
5.2.1.
Standard Non-Separating Mechanical Interface ..................................................................... 38
5.2.2.
Orbital Supplied Mechanical Interface Control Drawing ......................................................... 38
5.3.
Payload Electrical Interfaces....................................................................................................... 38
5.3.1.
Payload Umbilical Interfaces................................................................................................... 38
5.3.2.
Payload Interface Circuitry ...................................................................................................... 40
5.3.3.
Payload Battery Charging ....................................................................................................... 40
5.3.4.
Payload Command and Control .............................................................................................. 40
5.3.5.
Pyrotechnic Initiation Signals .................................................................................................. 40
5.3.6.
Payload Telemetry .................................................................................................................. 40
5.3.7.
Payload Separation Monitor Loopbacks ................................................................................. 41
5.3.8.
Telemetry Interfaces ............................................................................................................... 41
5.3.9.
Non-Standard Electrical Interfaces ......................................................................................... 41
5.3.10.
Electrical Launch Support Equipment................................................................................... 41
5.4.
Payload Design Constraints........................................................................................................ 41
5.4.1.
Payload Center of Mass Constraints ...................................................................................... 41
5.4.2.
Final Mass Properties Accuracy.............................................................................................. 41
5.4.3.
Pre-Launch Electrical Constraints........................................................................................... 42
5.4.4.
Payload EMI/EMC Constraints................................................................................................ 42
5.4.5.
Payload Dynamic Frequencies ............................................................................................... 42
5.4.6.
Payload Propellant Slosh ........................................................................................................ 42
5.4.7.
Payload-Supplied Separation Systems................................................................................... 42
5.4.8.
System Safety Constraints...................................................................................................... 42
6.
MISSION INTEGRATION.................................................................................................................. 43 6.1.
Mission Management Approach ................................................................................................. 43
6.1.1.
RSLP Mission Responsibilities................................................................................................ 43
6.1.2.
Orbital Mission Responsibilities .............................................................................................. 43
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6.2.
Mission Planning and Development ........................................................................................... 44
6.3.
Mission Integration Process........................................................................................................ 44
6.3.1.
Integration Meetings................................................................................................................ 44
6.3.2.
Mission Design Reviews (MDR).............................................................................................. 45
6.3.3.
Readiness Reviews................................................................................................................. 45
6.4.
Documentation............................................................................................................................ 46
6.4.1.
Customer-Provided Documentation ........................................................................................ 46
6.4.1.1.
Payload Questionnaire...................................................................................................... 46
6.4.1.2.
Payload Mass Properties .................................................................................................. 46
6.4.1.3.
Payload Finite Element Model .......................................................................................... 46
6.4.1.4.
Payload Thermal Model for Integrated Thermal Analysis................................................. 46
6.4.1.5.
Payload Drawings ............................................................................................................. 46
6.4.1.6.
Program Requirements Document (PRD) Mission Specific Annex Inputs ....................... 46
6.4.1.6.1. 6.5.
Launch Operations Requirements (OR) Inputs .......................................................... 47
Safety .......................................................................................................................................... 47
6.5.1.
System Safety Requirements.................................................................................................. 47
6.5.2.
System Safety Documentation................................................................................................ 47
7.
GROUND AND LAUNCH OPERATIONS ......................................................................................... 49 7.1.
Minotaur IV/Payload Integration Overview ................................................................................. 49
7.2.
Ground And Launch Operations ................................................................................................. 49
7.2.1.
Launch Vehicle Integration...................................................................................................... 49
7.2.1.1.
Planning and Documentation............................................................................................ 49
7.2.1.2.
GCA/Orion 38 Integration and Test Activities ................................................................... 49
7.2.1.3.
PK Motor Integration and Test Activities........................................................................... 49
7.2.1.4.
Mission Simulation Tests .................................................................................................. 49
7.2.1.5.
Booster Assembly Stacking/Launch Pad Preparation ...................................................... 51
7.2.2.
Payload Processing/Integration .............................................................................................. 51
7.2.2.1.
Payload Propellant Loading .............................................................................................. 52
7.2.2.2.
Final Vehicle Integration and Test .................................................................................... 52
7.3.
Launch Operations...................................................................................................................... 53
7.3.1. 8.
Launch Control Organization .................................................................................................. 53
OPTIONAL ENHANCED CAPABILITIES.......................................................................................... 55 8.1.
Mechanical Interface and Separation System Enhancements ................................................... 55
8.1.1.
Separation Systems ................................................................................................................ 55
8.1.2.
Additional Fairing Access Doors ............................................................................................. 55
8.1.3.
Payload Isolation System........................................................................................................ 55
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Performance Enhancements ...................................................................................................... 56
8.2. 8.2.1.
Insertion Accuracy................................................................................................................... 56
8.2.2.
Star 48 Stage 4 ....................................................................................................................... 57
8.3.
Environmental Control Options ................................................................................................... 58
8.3.1.
Conditioned Air........................................................................................................................ 58
8.3.2.
Nitrogen Purge ........................................................................................................................ 58
8.3.3.
Enhanced Contamination Control ........................................................................................... 58
8.3.3.1.
High Cleanliness Integration Environment (Class 10K or 100K) ...................................... 59
8.3.3.2.
Fairing Surface Cleanliness Options ................................................................................ 59
8.3.3.3.
High Cleanliness Fairing Environment.............................................................................. 59
8.3.4.
Launch Pad Environmental Control ........................................................................................ 59
8.3.4.1. 8.4.
Booster Temperature Control ........................................................................................... 59
Enhanced Telemetry Options ..................................................................................................... 60
8.4.1.
Enhanced Telemetry Bandwidth ............................................................................................. 60
8.4.2.
Enhanced Telemetry Instrumentation ..................................................................................... 60
8.4.3.
Navigation Data....................................................................................................................... 60
8.5.
Shared Launch Accommodations ............................................................................................... 60 LIST OF FIGURES
Figure 2-1. OSP-2 Peacekeeper Space Lift Vehicle.................................................................................... 3 Figure 2-2. Minotaur IV Expanded View ...................................................................................................... 4 Figure 2-3. Orion 38 Stage 4 Motor ............................................................................................................. 6 Figure 2-4. Stage 4 Structures..................................................................................................................... 6 Figure 2-5. Existing 92 in. Taurus Fairing, Handling Fixtures, and Processes will be used for Minotaur IV....................................................................................................... 7 Figure 2-6. Functional Block Diagram of LSE.............................................................................................. 9 Figure 2-7. Portable Launch Support Structure Provide Optional Support From Austere Sites................ 10 Figure 3-1. Minotaur IV Mission Profile...................................................................................................... 12 Figure 3-2. Minotaur IV Launch Site Options............................................................................................. 13 Figure 3-3. Minotaur IV Performance Curves for VAFB Launches (Metric Units) ..................................... 14 Figure 3-4. Minotaur IV Performance Curves for VAFB Launches (English Units) ................................... 15 Figure 3-5. Minotaur IV Performance Curves for CCAFS Launches (Metric Units) .................................. 16 Figure 3-6. Minotaur IV Performance Curves for CCAFS Launches (English Units) ................................ 17 Figure 3-7. Minotaur IV Performance Curves for Kodiak, Alaska Launches (Metric Units)....................... 18 Figure 3-8. Minotaur IV Performance Curves for Kodiak, Alaska Launches (English Units)..................... 19 Figure 3-9. Minotaur IV Performance Curves for Wallops, Virginia Launches (Metric Units).................... 20 Figure 3-10. Minotaur IV Performance Curves for Wallops, Virginia Launches (English Units)................ 21 Release 1.1
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Figure 3-11. Stage Impact Points for VAFB and CCAFS Launches.......................................................... 22 Figure 3-12. Minotaur IV Injection Accuracy .............................................................................................. 24 Figure 3-13. Typical Pre-Separation Payload Pointing and Spin Rate Accuracies .................................. 24 Figure 4-1. Phasing of Dynamic Loading Events....................................................................................... 25 Figure 4-2. Payload CG Net Transient Lateral Acceleration at Stage 2 Ignition with a Typical Separation System................................................................................................................... 27 Figure 4-3. Minotaur IV Nominal Maximum Axial Acceleration as a Function of Payload Mass ............... 27 Figure 4-4. Minotaur IV Payload Random Vibration Environment ............................................................. 29 Figure 4-5. Minotaur IV Payload Sine Vibration Environment ................................................................... 29 Figure 4-6. Minotaur IV Payload Acoustic Maximum Predicted Environment (MPE) ................................ 30 Figure 4-7. Minotaur IV Payload Shock Maximum Predicted Environment (MPE) – Launch Vehicle to Payload .................................................................................................................................... 31 Figure 4-8. Payload Shock Environment – Payload to Launch Vehicle .................................................... 32 Figure 4-9. Factors of Safety Payload Design and Test ........................................................................... 33 Figure 4-10. Recommended Payload Testing Requirements.................................................................... 33 Figure 4-11. Payload Thermal and Humidity Environment ........................................................................ 34 Figure 4-12. Minotaur IV Launch Vehicle RF Emitters and Receivers ...................................................... 36 Figure 5-1. Standard 92 in. Fairing Envelope ............................................................................................ 37 Figure 5-2. Standard, Non-separating Payload Mechanical Interface....................................................... 39 Figure 5-3. Payload Electrical Interface Block Diagram, With No Orbital Supplied Separation System .. 39 Figure 5-4. Payload 1:1 Umbilical Pin Outs ............................................................................................... 39 Figure 5-5. Minotaur IV Payload Electrical Interface Block Diagram ......................................................... 40 Figure 5-6. Payload Mass Properties Measurement Tolerance ................................................................ 42 Figure 6-1. Typical Integrated OSP Organizational Structure ................................................................... 43 Figure 6-2. Typical Mission Integration Schedule...................................................................................... 45 Figure 7-1. Hardware Flow – Factory to Launch Site ................................................................................ 50 Figure 7-2. SLV Processing Flow .............................................................................................................. 51 Figure 7-3. Minotaur IV Processing Flow................................................................................................... 52 Figure 7-4. Minotaur IV Upper Stack Assembly will be Vertically Integrated to Minotaur IV Booster Assembly in a Similar Manner to Taurus Upper Stack ............................................................ 53 Figure 8-1. 38-in. Separation System Option............................................................................................. 56 Figure 8-2. Soft Ride Payload Isolation System as Integrated on Minotaur LV......................................... 57 Figure 8-3. Hydrazine Auxiliary Propulsion System (HAPS) Used to Provide Insertion Accuracy............ 57 Figure 8-4. Orion 38 Stage 4 Motor can be Replaced with a Star-48 to Provide Increased Performance 58 Figure 8-5. Mobile Scaffolding for Environmental Control Demonstrated on Minotaur Missions .............. 59 Figure 8-6. Modular Minotaur IV Structural Design Easily Accommodates Multiple Payloads ................. 60
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Table of Contents LIST OF TABLES PAGE
TABLE 4-1. PAYLOAD CG PARAMETRIC DESIGN LIMIT LOADS ......................................................... 26
LIST OF APPENDICES A. PAYLOAD QUESTIONNAIRE..............................................................................................................A-1
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Minotaur IV User’s Guide ACS A/D AADC AC AFRL BER C/CAM CCAFS CDR CG CLA CLF CMP DPAF ECS EELV EGSE EMC EME EMI FLSA FTS GCA GFE GFP GPB GPS GTO HAPS HEPA HVAC ICD LCR LER LEV LOCC LRR LSE LSV LV MAC MAC
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Glossary
Attitude Control System Arm/Disarm Alaska Aerospace Development Corporation Air Conditioning Air Force Research Laboratory Bit Error Rate Collision/Contamination Avoidance Maneuver Cape Canaveral Air Force Station Critical Design Review center-of-gravity Coupled Loads Analysis Commercial Launch Facility Critical Measurements Program Dual Payload Attach Fitting Environmental Control System Evolved Expendable Launch Vehicle Electrical Ground Support Equipment Electromagnetic Compatibility Electromagnetic Environment Electromagnetic Interference Florida Spaceport Authority Flight Termination System Guidance and Control Assembly Government Furnished Equipment Government Furnished Property GPS Position Beacon Global Positioning System Geosynchronous Transfer Orbit Hydrazine Auxiliary Propulsion System High Efficiency Particulate Air Heating, Ventilation and Air Conditioning Interface Control Drawing Launch Control Room Launch Equipment Room Launch Equipment Van Launch Operations Control Center Launch Readiness Review Launch Support Equipment Launch Support Van Launch Vehicle Motor Adapter Cone Modified Adapter Cone
MACH MARS MDR MIWG MRD MRR NTW ODM OR OSP-2 PAF PAM PBCM PCF PCM PDR PID PK POC PPF PRD PWB QES RF RSLP RTS RWG SEA SEB SES SLV SRSS SSI START TPS TVAs TVC USAF VAFB VPF WFF WP
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Modular Avionics Control Hardware Mid-Atlantic Regional Spaceport Mission Design Review Mission Integration Working Groups Mission Requirements Document Mission Readiness Review Navy Theater Wide Ordnance Driver Modules Operations Requirements Orbital Suborbital Program ll Payload Attach Fitting Payload Adapter Module PK Booster Control Module Programmable Clock Filter Pulse Code Modulation Preliminary Design Review Proportional-Integral-Derivative Peacekeeper Point of Contact Payload Processing Facility Program Requirements Document Printed Wiring Board Quick Erect Scaffold Radio Frequency Rocket Systems Launch Program Range Tracking System Range Working Groups Statistical Energy Analysis Support Equipment Building Saab Ericson Space Space Launch Vehicle Softride for Small Satellites Spaceport Systems International Strategic Arms Reduction Treaty Thermal Protection System Thrust Vector Actuators Thrust Vector Control United States Air Force Vandenberg Air Force Base Vehicle Processing Facility Wallops Flight Facility Work Package
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Section 1.0 - Introduction
1. INTRODUCTION This User’s Guide is intended to familiarize payload mission planners with the capabilities of the Orbital Suborbital Program ll (OSP-2) Minotaur IV Space Launch Vehicle (SLV) launch service. The information provided in this user’s guide is for initial planning purposes only. Information for development/design is determined through mission specific engineering analyses. The results of these analyses are documented in a mission-specific Interface Control Document (ICD) for the payloader organization to use in their development/design process. This document provides an overview of the Minotaur IV system design and a description of the services provided to our customers. Minotaur IV offers a variety of enhanced options to allow the maximum flexibility in satisfying the objectives of single or multiple payloads.
By adopting the austere launch site concepts developed for Taurus, the Minotaur IV 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 Force Station (CCAFS), Wallops Flight Facility (WFF), and Kodiak Island, Alaska. This User’s Guide describes Minotaur IVunique integration and test approaches (including the typical operational timeline for payload integration with the Minotaur IV vehicle) and the ground support equipment that will be used to conduct Minotaur IV operations.
The primary mission of the Minotaur IV is to provide low cost, high reliability launch services to government-sponsored payloads. Minotaur IV accomplishes this using flight-proven components with a significant flight heritage such as surplus Peacekeeper boosters, the Taurus Fairing and Attitude Control System, and a mix of Minotaur I, Pegasus, Taurus, and other orbital standard avionics, all with a proven, successful track record. The philosophy of placing mission success as the highest priority is reflected in the success and accuracy of all OSP missions to date. The Minotaur IV launch vehicle system is composed of a flight vehicle and ground support equipment. Each element of the Minotaur IV system has been developed to simplify the mission design and payload integration process and to provide safe, reliable space launch services. This User’s Guide describes the basic elements of the Minotaur IV system as well as optional services that are available. In addition, this document provides general vehicle performance, defines payload accommodations and environments, and outlines the Minotaur IV mission integration process.
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Section 2.0 – Minotaur IV Launch Service
2. MINOTAUR IV LAUNCH SERVICE 2.1. Minotaur IV Launch System Overview The Minotaur IV (Figure 2-1) mission is to provide a cost effective, reliable and flexible means of placing satellites into orbit. The launch vehicle developer and manufacturer is Orbital, under the Orbital Suborbital Program 2 (OSP-2) contract for the U.S. Air Force. An overview of the system and available launch services is provided within this section, with specific elements covered in greater detail in the subsequent sections of this User’s Guide.
IV draws on the successful heritage of four launch vehicles: Orbital’s Pegasus, Taurus and Minotaur I space launch vehicles and the Peacekeeper system of the USAF. Minotaur IV’s avionics are derived from the Pegasus and Taurus systems, providing a combined total of more than 35 successful space launch missions. Orbital’s efforts have enhanced or updated Pegasus, Taurus and Minotaur avionics components to meet the payload-support requirements of the OSP-2 program. Combining these improved subsystems with the long successful history of the Peacekeeper boosters has resulted in a simple, robust, self-contained launch system to support government-sponsored small satellite launches. The Minotaur IV system also includes a complete set of transportable Launch Support Equipment (LSE) designed to allow Minotaur IV to be operated as a self-contained satellite delivery system. To accomplish this goal, the Electrical Ground Support Equipment (EGSE) has been developed to be portable and adaptable to varying levels of infrastructure. While the Minotaur IV system is capable of self-contained operation using portable vans to house the EGSE, it is typically launched from an established range where the EGSE can be housed in available, permanent structures or facilities.
Figure 2-1. OSP-2 Peacekeeper Space Lift Vehicle Minotaur IV has been designed to meet the needs of U.S. Government-sponsored customers at a lower cost than commercially available alternatives by the use of surplus Peacekeeper boosters. The requirements of the OSP-2 program stress system reliability, transportability, and operation from multiple launch sites. Minotaur
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The vehicle and LSE are designed to be capable of launch from any of the four commercial Spaceports (Alaska, California, Florida, and Mid-Atlantic), as well as from existing U.S. Government facilities at VAFB and CCAFS. The Launch Control Room (LCR) serves as the control center for conducting a Minotaur IV launch and includes consoles for Orbital, range safety, and limited customer personnel. Further description of the Launch Support Equipment is provided in Section 2.4. 2.2. Minotaur IV Launch Service The Minotaur IV Launch Service is provided through the combined efforts of the USAF and Orbital, along with associate contractors including
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Northrop Grumman and Commercial Spaceports. The primary customer interface will be with the USAF Space and Missile Systems Center, Detachment 12, Rocket Systems Launch Program, designated RSLP. Orbital is the launch vehicle provider. For brevity, this integrated team effort will be referred to as “OSP”. Where interfaces are directed toward a particular member of the team, they will be referred to directly (i.e., Orbital or RSLP). OSP provides all of the necessary hardware, software and services to integrate, test and launch a satellite into its prescribed orbit. In addition, OSP will complete all the required agreements, licenses and documentation to successfully conduct Minotaur IV operations. The Minotaur IV mission integration process completely identifies, documents, and verifies all spacecraft and mission requirements. This provides a solid basis for initiating and streamlining the integration process for future Minotaur IV customers. 2.3. Minotaur IV Launch Vehicle The Minotaur IV vehicle, shown in expanded view in Figure 2-2, is a four-stage, inertially guided, all solid propellant ground launched vehicle. Conservative design margins, state-ofthe-art structural systems, a modular avionics architecture, and simplified integration and test capability, yield a robust, highly reliable launch vehicle design. In addition, Minotaur IV payload accommodations and interfaces have been designed to satisfy a wide range of potential payload requirements. 2.3.1. Stage 1, 2 and 3 Booster Assemblies The first three stages of the Minotaur IV consists of the refurbished Government Furnished Equipment (GFE) Peacekeeper Stages 1, 2 and 3. These booster assemblies are used as provided by the Government, requiring no modification or additional components. They have extensive flight history, including 51 Peacekeeper launches and three Stage 1 launches on Taurus with no motor related failures. All three stages are solid-
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Figure 2-2. Minotaur IV Expanded View propellant motors and utilize a movable nozzle controlled by a Thrust Vector Actuator (TVA) system for three-axis attitude control. The first
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stage consists of a Thiokol motor that provides 500,000 lbf (2224 kN) of thrust. The second stage motor is an Aerojet motor with a moveable nozzle contoured with an extendable exit cone. It provides an average thrust of 275,000 lbf (1223 kN). The third stage is a Hercules motor that provides 65,000 lbf (289 kN) of thrust and also features an extendable exit cone similar to Stage 2. 2.3.2. Stage 4 Booster/Avionics Assembly The Minotaur IV Stage 4 motor is the Orion 38 design (Figure 2-3). This motor was originally developed for Orbital’s Pegasus program and is used on the Minotaur I launch vehicle, as well as other Orbital launch vehicles. Common design features, materials and production techniques are applied to the motors to maximize reliability and production efficiency. The Orion 38 motor provides the velocity needed for orbit insertion for the SLV, in the same manner as it is used on the Minotaur. This motor features state-of-the-art design and materials with a successful flight heritage. It is currently in production and is actively flying payloads into space, with 37 fully successful flights to date and one static test. Three substructures are utilized to accommodate the Orion 38 Stage 4 motor and attach it to the Stage 1-3 PK booster assembly. These are a Payload Adapter Module (PAM) with 62.01-inch diameter payload interface, a 38-inch diameter motor adapter cone and a GCA +3/4 interstage. These structures were adapted from similar Taurus hardware designs and are shown in Figure 2-4. 2.3.2.1. Avionics The Minotaur IV avionics system has heritage to the Minotaur I, OSP TLV, as well as Pegasus and Taurus designs. The flight computer is a 32bit multiprocessor architecture. It provides communication with vehicle subsystems, the LSE, and the payload, if required, utilizing standard RS422 serial links and discrete I/O. The avionics system design incorporates Orbital’s innovative,
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flight-proven Modular Avionics Control Hardware (MACH). Standardized, function-specific modules are combined in stacks to meet vehicle-specific requirements. The functional modules from which the MACH stacks are created include power transfer, ordnance initiation, booster interface, communication, and telemetry processing. Orbital has designed, tested, and flown a variety of MACH modules, which provide an array of functional capability and flexibility. MACH has exhibited 100% reliability on all flights including Minotaur and TLV flights and several of Orbital’s suborbital launch vehicles including Navy Theater Wide (NTW), Critical Measurements Program (CMP), and STORM. 2.3.2.2. Attitude Control Systems The Minotaur IV Attitude Control System (ACS) provides attitude control throughout boosted flight and coast phases. Stages 1, 2 and 3 utilize the PK Thrust Vector Control (TVC) systems. The PK Booster Control Module (PBCM) links the flight computer actuator commands to the individual Stage 1, 2, and 3 Thrust Vector Actuators (TVAs). Stage 4 utilizes the same TVC system used by the Pegasus, Taurus and Minotaur vehicles which combines single-nozzle electromechanical TVC for pitch and yaw control with a three-axis, cold-gas attitude control system resident in the avionics section providing roll control. Attitude control is achieved using a three-axis autopilot that employs Proportional-IntegralDerivative (PID) control. Stages 1, 2 and 3 fly a pre-programmed attitude profile based on trajectory design and optimization. Stage 4 uses a set of pre-programmed orbital parameters to place the vehicle on a trajectory toward the intended insertion apse. The extended coast between Stages 3 and 4 is used to orient the vehicle to the appropriate attitude for Stage 4 ignition based upon a set of pre-programmed orbital parameters and the measured performance of the first three stages. Stage 4 utilizes energy management to place the vehicle into the proper orbit. After the
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Figure 2-3. Orion 38 Stage 4 Motor
Figure 2-4. Stage 4 Structures
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final boost phase, the three-axis cold-gas attitude control system is used to orient the vehicle for spacecraft separation, contamination and collision avoidance and downrange downlink maneuvers. The autopilot design is modular, so additional payload requirements such as rate control or celestial pointing can be accommodated with minimal additional development. 2.3.2.3. Telemetry Subsystem The Minotaur IV telemetry subsystem provides real-time health and status data of the vehicle avionics system, as well as key information regarding the position, performance and environment of the Minotaur IV vehicle. This data may be used by Orbital and the range safety personnel to evaluate system performance. The Minotaur IV baseline telemetry subsystem provides a number of dedicated payload discrete (bi-level) and analog telemetry monitors through dedicated channels in the SLV encoder. The Minotaur IV telemetry system provides a baseline 1 Mbps data rate (both payload and Minotaur IV telemetry). However, the output data rate is selectable from 2.441 kbps to 10 Mbps to allow flexibility in supporting evolving mission requirements, as limited by link margin and Bit Error Rate (BER) requirements. The telemetry subsystem nominally utilizes Pulse Code Modulation (PCM) with a RNRZ-L format. However other types of data formats, including NRZ-L, S, M, and Bi-phase may be implemented if required, in order to accommodate launch range limitations.
control during ground handling, integration operations and flight. The fairing is a bi-conic design made of graphite/epoxy face sheets with an aluminum honeycomb core. The fairing provides for low payload contamination through prudent design and selection of low contamination materials and processes. Acoustic blankets and internal air conditioning ducts are available to provide more benign payload environments. Air conditioning will keep the payload environment to a specified temperature between 60 to 120 °F dependent upon requirements. The two halves of the fairing are structurally joined along their longitudinal interface using Orbital’s low contamination frangible joint system. An additional circumferential frangible joint at the base of the fairing supports the fairing loads. At separation, a gas pressurization system is activated to pressurize the fairing deployment thrusters. The fairing halves then rotate about external hinges that control the fairing deployment to ensure that payload and launch vehicle clearances are maintained. All elements of the deployment system have been demonstrated through test to comply with stringent contamination requirements. Options for payload access doors and enhanced cleanliness are available. Further
Minotaur IV telemetry is subject to the provisions of the Strategic Arms Reduction Treaty (START). START treaty provisions require that certain Minotaur IV telemetry be unencrypted and provided to the START treaty office for dissemination to the signatories of the treaty. 2.3.2.4. Payload Fairing Orbital’s flight-proven Taurus 92-inch diameter payload fairing (Figure 2-5) is used to encapsulate the payload, provide protection and contamination
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Figure 2-5. Existing 92 in. Taurus Fairing, Handling Fixtures, and Processes will be used for Minotaur IV
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details on the baseline fairing are included in Section 5.1. With the addition of a structural adapter, the fairing can accommodate multiple payloads. This feature, described in more detail in Section 8.5 of this User’s Guide, permits two or more smaller payloads to share the cost of a Minotaur launch, resulting in a lower launch cost for each as compared to other launch options. OSP has access to several Multiple Payload Adaptor (MPA) designs that allow for a cost sharing benefit to programs with excess payload and/or mass capability. 2.4. Launch Support Equipment The Minotaur IV LSE is designed to be readily adaptable to varying launch site configurations with minimal unique infrastructure required. The EGSE consists of readily transportable consoles that can be housed in various facility configurations depending on the launch site infrastructure. The EGSE is composed of three primary functional elements: Launch Control, Vehicle Interface, and Telemetry Data Reduction. The Launch Control consoles are located in a Launch Control Room (LCR), or mobile launch equipment van depending on available launch site accommodations. The Vehicle Interface EGSE is located in a permanent structure, typically called a Support Equipment Building (SEB) or Launch Equipment Room (LER). Fiber optic connections from the Launch Control to the Vehicle Interface consoles are used for efficient, high bandwidth communications and eliminates the need for copper wire. The Vehicle Interface racks provide the junction from fiber optic cables to the copper cabling interfacing with the vehicle. Figure 2-6 depicts the functional block diagram of the LSE. The LCR serves as the control center during the launch countdown. The number of personnel that can be accommodated are dependent on the launch site facilities. At a minimum, the LCR will accommodate Orbital personnel controlling the vehicle, two Range Safety representatives (ground and flight safety), and the Air Force Mission Manager. Mission-unique, customer-supplied
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payload consoles and equipment can be supported in the LCR and SEB, within the constraints of the launch site facilities or temporary structure facilities. Interface to the payload through the Minotaur IV payload umbilicals and landlines provides the capability for direct monitoring of payload functions. Payload personnel accommodations will be handled on a mission-specific basis. 2.4.1. Transportable LSE Shelters In order to perform mission operations from alternative, austere launch sites, Orbital can provide self contained, transportable shelters for the Launch Support Van (LSV) and Launch Equipment Van (LEV) as an unpriced option. These shelters are the same approach and design used on all six of Orbital’s Taurus launches (Figure 2-7). The OSP-2 Ground Support Consoles have been intentionally made modular and portable to allow their use in these accommodations. The LSV consists of a shelter which is located at a Range Safety-approved man-safe distance from the launch site. The LSV contains the vehicle control and telemetry monitor consoles described in Section 2.4. Sufficient space is available for additional equipment racks depending on Government and/or Payload requirements. The LEV consists of an 8 foot x 20 foot shelter which is located near the launch stool and is unmanned during launch. This shelter contains the vehicle interface racks described in Section 2.4. The LEV has sufficient room available for payload power supplies and interface electronics. Both shelters are designed for shipping and transportation with exterior tiedown and anchor locations used to facilitate the loading and unloading operations. The shelters can be delivered to any level location and be set up within hours. The transportable support console design also allows for the LSE to be moved into a fixed blockhouse and LER if required.
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Figure 2-6. Functional Block Diagram of LSE
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Figure 2-7. Portable Launch Support Structure Provide Optional Support From Austere Sites
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3. GENERAL PERFORMANCE
3.2.2. Eastern Launch Sites For Easterly launch azimuths to achieve orbital inclinations between 28.5° and 60°, Minotaur IV can be launched from facilities at Cape Canaveral, FL or Wallops Island, VA. Launches from Florida will nominally use the Florida Spaceport Authority (FLSA) launch facilities at LC-46 on CCAFS. These will be typically for inclinations from 28.5° to 40°, although inclinations above 35° may have reduced performance due to the need for a trajectory dogleg. Inclinations below 28.5° are feasible, albeit with doglegs and altitude constraints due to range safety considerations. The Mid-Atlantic Regional Spaceport (MARS) facilities at the WFF may be used for inclinations from 30° to 60°. Some inclinations may have reduced performance due to range safety considerations and will need to be evaluated on a case-by-case missionspecific basis.
3.1. Mission Profiles Minotaur IV can attain a range of posigrade and retrograde inclinations through the choice of launch sites made available by the readily adaptable nature of the Minotaur IV launch system. A typical mission profile to a sunsynchronous orbit is shown in Figure 3-1. All performance parameters presented herein are typical for most expected payloads. However, performance may vary depending on unique payload or mission characteristics. Specific requirements for a particular mission should be coordinated with OSP. Once a mission is formally initiated, the requirements will be documented in the Mission Requirements Document (MRD). Further detail will be captured in the Payload-toLaunch Vehicle Interface Control Drawing (ICD). 3.2. Launch Sites Depending on the specific mission and range safety requirements, Minotaur IV can operate from several East and West Coast launch sites, illustrated in Figure 3-2. Specific performance parameters are presented in Section 3.3. Facilities used for OSP Minotaur and Taurus launches are generally compatible with Minotaur IV operations. 3.2.1. Western Launch Sites For missions requiring high inclination orbits (greater than 60°), launches can be conducted from facilities at VAFB or Kodiak Island, AK. Both facilities can accommodate inclinations from 60° to 120°, although inclinations below 72° from VAFB require an out-of-plane dogleg, thereby reducing payload capability. As with the initial Minotaur missions, Minotaur IV can be launched from the California Spaceport facility, Space Launch Complex 8 (SLC-8) operated by Spaceport Systems International (SSI), on South VAFB. The launch facility at Kodiak Island, operated by the Alaska Aerospace Development Corporation (AADC) has been used for both orbital and suborbital launches.
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3.2.3. Alternate Launch Sites Other launch facilities can be readily used given the flexibility designed into the Minotaur IV vehicle, ground support equipment, and the various interfaces. Orbital has experience launching vehicles from a variety of sites around the world. To meet the requirements of performing mission operations from alternative, austere launch sites, Orbital can provide self contained, transportable shelters as described in section 2.4.1. 3.3. Performance Capability Minotaur IV performance curves for circular and elliptical orbits of various altitudes and inclinations are detailed in Figure 3-3 through Figure 3-10 for launches from all four Spaceports in both metric and English units. These performance curves provide the total mass above the standard, non-separating interface. The mass of any Payload Attach Fitting (PAF) or separation system is to be accounted for in the payload mass allocation. Figure 3-11 illustrates the stage vacuum impact points for launch trajectories from VAFB and CCAFS.
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Figure 3-1. Minotaur IV Mission Profile
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Figure 3-2. Minotaur IV Launch Site Options
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Figure 3-3. Minotaur IV Performance Curves for VAFB Launches (Metric Units)
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Figure 3-4. Minotaur IV Performance Curves for VAFB Launches (English Units)
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Figure 3-5. Minotaur IV Performance Curves for CCAFS Launches (Metric Units)
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Figure 3-6. Minotaur IV Performance Curves for CCAFS Launches (English Units)
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Figure 3-7. Minotaur IV Performance Curves for Kodiak, Alaska Launches (Metric Units)
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Figure 3-8. Minotaur IV Performance Curves for Kodiak, Alaska Launches (English Units)
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Figure 3-9. Minotaur IV Performance Curves for Wallops, Virginia Launches (Metric Units)
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Figure 3-10. Minotaur IV Performance Curves for Wallops, Virginia Launches (English Units)
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Figure 3-11. Stage Impact Points for VAFB and CCAFS Launches
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3.4. Injection Accuracy Minotaur IV injection accuracy is summarized in Figure 3-12. Better accuracy can be provided dependent on specific mission characteristics. For example, heavier payloads will typically have better insertion accuracy, as will higher orbits. An enhanced option for increased insertion accuracy is also available (Section 8.2.1). It utilizes the flight-proven Hydrazine Auxiliary Propulsion System (HAPS) developed on the Pegasus program. 3.5. Payload Deployment Following orbit insertion, the Minotaur IV Stage 4 avionics subsystem can execute a series of ACS maneuvers to provide the desired initial payload attitude prior to separation. This capability may also be used to incrementally reorient Stage 4 for the deployment of multiple spacecraft with independent attitude requirements. Either an inertially-fixed or spin-stabilized attitude may be specified by the customer. The maximum spin rate for a specific mission depends upon the spin axis moment of inertia of the payload and the amount of ACS propellant needed for other attitude maneuvers. Figure 3-13 provides the typical payload pointing and spin rate accuracies. 3.6. Payload Separation Payload separation dynamics are highly dependent on the mass properties of the payload and the particular separation system utilized. The primary parameters to be considered are payload tip-off and the overall separation velocity.
Separation velocities are driven by the need to prevent recontact between the payload and the Minotaur IV 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 IV 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 IV 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 “crabwalk” maneuver to impart a small amount of delta velocity, increasing the separation between the payload and the fourth stage of the Minotaur IV. Following the completion of the C/CAM maneuver as described above and any remaining maneuvers, such as downlinking of delayed telemetry data (per START treaty provisions), the ACS valves are opened and the remaining ACS nitrogen propellant is expelled.
Payload tip-off refers to the angular velocity imparted to the payload upon separation due to payload center-of-gravity (CG) offsets and an uneven distribution of torques and forces. If an optional Orbital-supplied Marmon Clamp-band separation system is used, payload tip-off rates are generally under 5°/sec per axis. Separation system options are discussed further in Section 8.1.1. Orbital performs a mission-specific tip-off analysis for each payload.
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Error Type
Section 3.0 – General Performance
Tolerance (Worst Case)
Error Source
Altitude (Insertion Apse)
±10 nmi (18.5 km)
Stage 4 motor performance uncertainty and guidance algorithm uncertainty
Altitude (Non-Insertion Apse)
±50 nmi (92.6 km)
Stage 4 motor performance and guidance algorithm uncertainty and navigation (INS) error
Altitude (Mean)
±30 nmi (55.6 km)
Stage 4 motor performance and guidance algorithm uncertainty and navigation (INS) error
Inclination
±0.2°
Guidance algorithm uncertainty and navigation (INS) error
Figure 3-12. Minotaur IV Injection Accuracy
Figure 3-13. Typical Pre-Separation Payload Pointing and Spin Rate Accuracies
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4. PAYLOAD ENVIRONMENT
events such as motor ignition, stage separation, and transonic crossover.
CAUTION The environmental design and test criteria presented have been derived using measured data obtained from previous Taurus, Minotaur I, and Peacekeeper missions, motor static fire tests, other system development tests and analyses. The predicted levels presented are intended to be representative of a standard mission. Satellite mass, geometry and structural components vary greatly and will result in significant differences from mission to mission.
The predicted environments provided in this user's guide are for initial planning purposes only. Environments presented here bound typical mission parameters, but should not be used in lieu of mission-specific analyses. Mission-specific levels are provided as a standard service and documented or referenced in the mission ICD. This section provides details of the predicted environmental conditions that the payload will experience during Minotaur IV ground operations, powered flight, and launch system on-orbit operations. The predicted environments provided in this user’s guide are for initial planning purposes only. Minotaur IV ground operations include payload integration and encapsulation within the fairing, subsequent transportation to the launch site and final vehicle integration activities. Powered flight begins at Stage 1 ignition and ends at Stage 4 burnout. Minotaur IV on-orbit operations begin after Stage 4 burnout and end following payload separation. To more accurately define simultaneous loading and environmental conditions, the powered flight portion of the mission is further subdivided into smaller time segments bounded by critical, transient flight
Dynamic loading events that occur throughout various portions of the flight include steady-state acceleration, transient low frequency acceleration, acoustic impingement, random vibration, and pyrotechnic shock events. Figure 4-1 identifies the time phasing of these dynamic loading events and environments and their significance. Pyroshock events are not indicated in this figure, as they do not occur simultaneous with any other significant dynamic loading events. In addition, dynamic loading associated with S4 ignition is insignificant. 4.1. Steady State and Transient Acceleration Loads Design limit load factors due to the combined effects of steady state and low frequency transient accelerations are largely governed by payload characteristics. A mission-specific Coupled Loads Analysis (CLA) will be performed, with customer
Figure 4-1. Phasing of Dynamic Loading Events Release 1.1
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provided finite element models of the payload, in order to provide precise load predictions. Results will be referenced in the mission specific ICD. For preliminary design purposes, Orbital can provide initial Center-of-Gravity (CG) netloads given a payload’s mass properties, CG location and bending frequencies. Design limit loads due to both transient and steady-state accelerations are presented in Table 4-1 for select payload masses.
vehicle model. A Monte-Carlo analysis is performed to determine variations in vehicle acceleration due to changes in winds, motor performance and aerodynamics. The steady-state accelerations must be added to transient accelerations from the CLA as indicated in Figure 4-1 to determine the total payload acceleration. Steady-state accelerations are typically 8-Gs axial and 0.5-Gs lateral.
4.1.1. Transient Loads Transient loads account for approximately 30% of the total vehicle load with the remainder due to steady wind loads. Typical acceleration values at the payload interface are 2-Gs lateral and 4-Gs axial depending on the load case. Transient lateral accelerations at Stage 2 ignition are defined as a function of payload mass in Figure 4-2. Preliminary and final CLAs will be performed for each Minotaur IV payload. The payload finite element model is coupled to the vehicle model. Forcing functions have been developed for different flight events or load cases. Load cases include liftoff, transonic, max q and stage and shroud separation.
During powered flight, the maximum steady state accelerations are dependent on the payload mass. The maximum level occurs during Stage 3 burn. Figure 4-3 depicts the axial acceleration at burnout for Stages 3 and 4 as a function of payload mass.
4.1.2. Steady-State Acceleration The steady-state vehicle accelerations are determined from the vehicle rigid body analysis. Drag, wind and motor thrust are applied to a
4.2.
Payload Vibration Environment The Minotaur IV payload vibration environments are low frequency random and sinusoidal vibrations created by the solid rocket motor combustion processes and transmitted through the launch vehicle structure. Additionally, higher frequency aeroacoustics energy is created by air flow over the surface of the vehicle. Some of this aeroacoustic energy is transmitted via the launch vehicle structure to the payload. However, a majority of the aeroacoustic energy is transmitted to the payload directly as acoustic energy through the fairing.
TABLE 4-1. PAYLOAD CG PARAMETRIC DESIGN LIMIT LOADS Payload Mass 1600 lbm (725.7 kg) Axial (G) max/min Liftoff
3.83/0.27
2400 lb (1089 kg)
Lateral (G)
Axial (G)
0.62
3.93/0.15
max/min
3200 lb (1452 kg)
Lateral (G)
Axial (G)
0.46
3.90/0.16
max/min
4000lb (8141 kg)
Lateral (G)
Axial (G)
0.41
4.01/0.12
max/min
Lateral (G) 0.37
Pre-Transonic Resonant Burn
5.05/0.83
0.02
3.46/2.29
0.00
3.44/2.31
0.00
3.73/2.08
0.00
Transonic
5.13/1.52
1.23
3.95/2.71
0.97
3.89/2.75
0.90
4.07/2.49
0.89
Supersonic
3.41/3.40
1.96
3.38/3.38
1.61
3.36/3.36
1.40
3.34/3.34
1.26
Stage 2 Ignition
3.93/-0.35
4.05
3.95/-0.03
2.89
3.83/-0.02
2.74
3.66/0.01
2.13
Stage 3 Ignition
6.79/0.00
0.78
6.45/0.00
0.59
6.22/0.00
0.49
5.90/0.00
0.40
Stage 3 Burnout
See Figure 4-3
TBS
See Figure 4-3
TBS
See Figure 4-3
TBS
See Figure 4-3
TBS
Stage 4 Burnout
See Figure 4-3
TBS
See Figure 4-3
TBS
See Figure 4-3
TBS
See Figure 4-3
TBS
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Figure 4-2. Payload CG Net Transient Lateral Acceleration at Stage 2 Ignition with a Typical Separation System
Figure 4-3. Minotaur IV Nominal Maximum Axial Acceleration as a Function of Payload Mass
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4.2.1. Random Vibration Random vibration environments are produced both by the structurally transmitted vibrations through mechanical payload interface as well as from the acoustic energy directly through the fairing (Figure 4-4). However, the higher frequency aeroacoustic-induced random vibration levels are hard to accurately replicate for spacecraft of the Minotaur IV class. Testing input levels are typically hard to accurately specify because the response at the LV-to-spacecraft interface is strongly dependent on the unique spacecraft dynamics, including its response to the acoustic field. Therefore, structure-born random vibration environments are only defined herein for frequencies up to 250 Hz with the recommendation that high frequency environments be based on the acoustic levels defined in Section 4.3. The acoustic environment is defined starting at 20 Hz, allowing both environments to be evaluated in the overlapping region. The structure-borne low frequency random vibration environment for Minotaur IV is a flat 0.002 g2/Hz from 20 to 250 Hz, producing an overall level of 0.68 gRMS (Figure 4-4). This level is based on measured data from multiple applicable Orbital launch vehicles. Vibration produced by the PK Stage 1 motor was analyzed based on data from three Taurus flights, which utilized the PK Stage 1 motor. The random vibration levels also cover the Orion 38 motor burn observed on past Minotaur, Pegasus, and Taurus flights. Detailed evaluation of the Stage 2 and 3 motor burn is still pending, but prior experience indicates that their levels will be relatively benign for the payload. The Minotaur IV structure-borne vibration environment is enveloped by the MILSTD-1540E levels. OSP recommends that the 1540E level be used as guidance for minimum levels for payload testing. 4.2.2. Sine Vibration A known resonant burn characteristic of the PK Stage 1 motor creates a sine vibration requirement of Minotaur IV payloads. Orbital has Release 1.1
reviewed test data from more than 30 static fire tests and three Taurus flights. The sine vibration level depends on payload weight and stiffness. Orbital has developed forcing function to simulate the Stage 1 motor resonance. The forcing functions are used in a NASTRAN simulation to predict peak sine vibration levels at different vehicle locations. The simulation has been validated using flight data. The sine vibration varies between 45 and 75 Hz. The NASTRAN simulation of the PK Stage 1 resonant burn is analyzed for each payload as part of the CLA discussed in Section 4.1. The results of this mission-specific CLA are then used to refine the SLV vehicle and payload sine vibration levels. A parametric analysis of hypothetical payloads has been conducted to provide guidance on expected acceleration levels. The resulting levels are shown in Figure 4-5. 4.3.
Payload Acoustic Environment The acoustic environments to which the spacecraft will be exposed have been defined based on measured acoustic data from previous Taurus flights which utilized the Peacekeeper Stage 1 motor and same 92 in. fairing. The data was adjusted to account for differences in vehicle trajectories and a 3 dB uncertainty was added per MIL-STD-1540 guidelines. The resulting acoustic level is shown in Figure 4-6. 4.4.
Payload Shock Environment The maximum shock response spectrum at the base of the payload from all launch vehicle events should not exceed the flight limit levels in Figure 4-7 (clamp-band separation system). Lower separation shock levels can be achieved by the use of a Lightband separation system provided by Planetary Systems Incorporated. The resulting levels are also shown in Figure 4-7. For missions that do not utilize an Orbitalsupplied payload separation system, the shock response spectrum at the base of the payload from vehicle events should not exceed the levels in Figure 4-7 (non-separating shock).
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Figure 4-4. Minotaur IV Payload Random Vibration Environment
Figure 4-5. Minotaur IV Payload Sine Vibration Environment Release 1.1
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Figure 4-6. Minotaur IV Payload Acoustic Maximum Predicted Environment (MPE)
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Figure 4-7. Minotaur IV Payload Shock Maximum Predicted Environment (MPE) – Launch Vehicle to Payload
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If the payload employs a non-Orbital separation system, then the shock delivered from the separation systems to the Stage 4 vehicle interface must not exceed the limit level characterized in Figure 4-8. Shock above this level could require requalification of components or an acceptance of risk by the Rocket Systems Launch Program (RSLP). 4.5.
Payload Structural Integrity and Environments Verification The primary support structure for the spacecraft must possess sufficient strength, rigidity, and other characteristics required to survive the critical loading conditions that exist within the envelope of handling and mission requirements, including worst-case predicted ground, flight, and post-boost loads. It must survive those conditions in a manner that assures safety and that does not reduce the mission success probability. Spacecraft design loads are defined as follows:
a. Design Limit Load — The maximum predicted ground-based, powered flight or on-orbit load, including all uncertainties. b. Design Yield Load — The Design Limit Load multiplied by the recommended Yield Factor of Safety (YFS). The payload structure must have sufficient strength to withstand simultaneously the yield loads, applied temperature, and other accompanying environmental phenomena for each design condition without experiencing detrimental yielding or permanent deformation. c. Design Ultimate Load — The Design Limit Load multiplied by the recommended Ultimate Factor of Safety (UFS). The payload structure must have sufficient strength to withstand simultaneously the ultimate loads, applied temperature, and other accompanying environmental phenomena without experiencing any fracture or other failure mode of the structure.
Figure 4-8. Payload Shock Environment – Payload to Launch Vehicle Release 1.1
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4.5.1. Recommended Payload Testing and Analysis Sufficient payload testing and/or analysis must be performed to ensure the safety of ground crews and to ensure mission success. The payload structural design should comply with the testing and design factors of safety in Figure 4-9. Vibration testing should be based on the standard margins defined in Figure 4-10. At a minimum, it is recommended that the following tests be performed: a. Structural Integrity — Static loads, sine vibration, or other tests should be performed that combine to encompass the acceleration load environment presented in Section 4.1. b. Random Vibration — The flight level environment and recommended test level is defined in Section 4.2.1. c. Acoustics — Full scale acoustic testing is recommended to verify higher frequency dynamics of the spacecraft are not adversely affected. The acoustic levels are defined in Figure 4-6. d. Shock – The payload separation event should be simulated to verify the spacecraft is not adversely effected. Shock levels are defined in Section 4.4. The payload organization must provide OSP with a list of the tests and test levels to which the payload was subjected prior to payload arrival at the integration facility. 4.6. Thermal and Humidity Environments The thermal and humidity environment to which the payload may be exposed during vehicle processing and pad operations are defined in the sections that follow and listed in Figure 4-11. 4.6.1. Ground Operations Upon encapsulation within the fairing and for the remainder of ground operations, the payload environment will be maintained by a Heating, Ventilation and Air Conditioning (HVAC) system.
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Figure 4-9. Factors of Safety Payload Design and Test
Figure 4-10. Recommended Payload Testing Requirements
The HVAC provides conditioned air to the payload in the PPF after fairing integration. The HVAC is used at the launch pad after vehicle stacking operations. Air Conditioning (AC) is not provided during transport or lifting operations without the enhanced option that includes High Efficiency Particulate Air (HEPA) filtration. The conditioned air enters the fairing at a location forward of the payload, exits aft of the payload and is provided up to 5 minutes prior to launch. Baffles are provided at the air conditioning inlet to reduce impingement velocities on the payload if required.
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Figure 4-11. Payload Thermal and Humidity Environment Fairing inlet conditions are selected by the customer, and are bounded as follows: a. Dry Bulb Temperature: 55 to 85 °F (13 to 29 °C) controllable to ±4 °F (±2 °C) of setpoint b. Dew Point Temperature: 38 to 62 °F (3 to 17 °C) c. Relative Humidity: determined by drybulb and dewpoint temperature selections and generally controlled to within ±15%. Relative humidity is bound by the psychrometric chart and will be controlled such that the dew point within the fairing is never reached. d. Maximum Flow: 500 cfm
than 0.1. This temperature limit envelopes the maximum temperature of any component inside the payload fairing with a view factor to the payload with the exception of the Stage 4 motor. The maximum Stage 4 motor surface temperature exposed to the payload will not exceed 350 °F (177 °C), assuming no shielding between the aft end of the payload and the forward dome of the motor assembly. The Payload Adapter Module (PAM), used with the fairing to provide encapsulation of the payload during ground processing, provides some level of shielding between the payload and Stage 4 motor. Whether this temperature is attained prior to payload separation is dependent upon mission timeline.
4.6.2. Powered Flight The maximum fairing inside wall temperature will be maintained at less than 200 °F (93 °C), with an emissivity of 0.92 in the region of the payload. As a non-standard service, a low emissivity coating can be applied to reduce emissivity to less
The fairing peak vent rate is typically less than 0.6 psi/sec. Fairing deployment will be initiated at a time in flight that the maximum dynamic pressure is less than 0.01 psf or the maximum free molecular heating rate is less than 0.1 BTU/ft2/sec, as required by the payload.
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4.6.3. Nitrogen Purge (Non-Standard Service) If required for spot cooling of a payload component, Orbital will provide GN2 flow to localized regions in the fairing as a non-standard service. This option is discussed in more detail in Section 8.3.2. 4.7. Payload Contamination Control All payload integration procedures, and Orbital’s contamination control program have been designed to minimize the payload’s exposure to contamination from the time the payload arrives at the payload processing facility through orbit insertion and separation. The payload is fully encapsulated within the fairing and Payload Adapter Module (PAM) at the payload processing facility, assuring the payload environment stays clean in a Class 100,000 environment. All SLV assemblies that affect cleanliness within the encapsulated payload volume include the fairing and the payload cone assembly. These assemblies are cleaned such that there is no particulate or non-particulate matter visible to the normal unaided eye when inspected from 2 to 4 feet under 50 ft-candle incident light (Visibly Clean Level II). After encapsulation, the fairing envelope is either sealed or maintained with a positive pressure, Class 100,000 continuous purge of conditioned air. If required, the payload can be provided with enhanced contamination control as an enhanced option, providing a Class 10,000 environment, low outgassing, and Visibly Clean Plus Ultraviolet cleanliness (see Section 8.3.3). Provisions exist in the fairing design accommodate dry nitrogen purge as has been demonstrated on the Taurus application. With the enhanced contamination control option, the Orbital-supplied elements will be cleaned and controlled to support a Class 10,000 clean room environment, as defined in Federal Standard 209. This includes limiting volatile hydrocarbons to maintain hydrocarbon content at less than 15 ppm and humidity between 35 to 60
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percent. Since the payload processing will be at a GFP facility, it is assumed the Class 10,000 clean room environment also adhering to these levels of control will be provided by that facility. Also with the enhanced contamination control option, Orbital provides an Environmental Control System (ECS) from payload encapsulation through vehicle lift-off. The ECS continuously purges the fairing volume with clean filtered air. Orbital’s ECS incorporates a HEPA filter unit to provide FED-STD-209 Class M5.5 (10,000) air. Orbital monitors the supply air for particulate matter via a probe installed upstream of the fairing inlet duct prior to connecting the air source to the payload fairing. 4.8. Payload Electromagnetic Environment The payload Electromagnetic Environment (EME) results from two categories of emitters: 1) Minotaur IV onboard antennas and, 2) Range radar. All power, control and signal lines inside the payload fairing are shielded and properly terminated to minimize the potential for Electromagnetic Interference (EMI). The Minotaur IV payload fairing is Radio Frequency (RF) opaque, which shields the payload from external RF signals while the payload is encapsulated. Figure 4-12 lists the frequencies and maximum radiated signal levels from vehicle antennas that are located near the payload during ground operations and powered flight. The specific EME experienced by the payload during ground processing at the PPF and the launch site will depend somewhat on the specific facilities that are utilized as well as operational details. However, typically the field strengths experienced by the payload during ground processing with the fairing in place are controlled procedurally and will be less than 2 V/m from continuous sources and less than 10 V/m from pulse sources. The highest EME during powered flight is created by the CBand transponder transmission, which results in
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peak levels at the payload interface plane of 88 V/m at 5765 MHz (based on Taurus). Range transmitters are typically controlled to provide a field strength of 10 V/m or less inside the fairing.
This EME should be compared to the payload’s RF susceptibility levels (MIL-STD-461, RS03) to define margin.
Figure 4-12. Minotaur IV Launch Vehicle RF Emitters and Receivers
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5. PAYLOAD INTERFACES This section describes the available mechanical, electrical and Launch Support Equipment (LSE) interfaces between the Minotaur IV launch vehicle and the payload. 5.1. Payload Fairing Orbital’s flight-proven Taurus 92-inch diameter payload fairing is used to encapsulate the payload, provide protection and contamination control during ground handling, integration operations and flight. The fairing is a bi-conic design made of graphite/epoxy face sheets with aluminum honeycomb core. The two halves of the fairing are structurally joined along their longitudinal interface using Orbital’s low contamination frangible joint system. An additional circumferential frangible joint at the base of the fairing supports the fairing loads. At separation, a gas pressurization system is activated to pressurize the fairing deployment thrusters. The fairing halves then rotate about external hinges that control the fairing deployment to ensure that payload and launch vehicle clearances are maintained. All elements of the deployment system have been demonstrated through test to comply with stringent contamination requirements.
payload deflection must be given with the Finite Element Model to evaluate payload dynamic deflection with the Coupled Loads Analysis (CLA). The payload contractor should assume that the interface plane is rigid; Orbital has accounted for deflections of the interface plane. The CLA will provide final verification that the payload does not violate the dynamic envelope. 5.1.2. Payload Access Door Orbital provides one 18 in. by 24 in. (45.7 cm by 61.0 cm) payload fairing access door to provide access to the payload after fairing mate. The door can be positioned according to payload requirements within the cylindrical section of the fairing, providing access to the payload without having to remove any portion of the fairing or break electrical connections. The specific location is defined and controlled in the payload ICD. Additional access doors can readily be provided as an enhanced option (see Section 8.1.2). 5.2. Payload Mechanical Interface and Separation System Minotaur IV provides for a standard nonseparating payload interface and an optional Orbital-provided payload separation system.
5.1.1. Payload Dynamic Design Envelope The fairing drawing in Figure 5-1 shows the maximum dynamic envelope available for the payload during powered flight. The dynamic envelope shown account for fairing and vehicle structural deflections only. The payload contractor must consider deflections due to spacecraft design and manufacturing tolerance stack-up within the dynamic envelope. Proposed payload dynamic envelope violations must be approved by OSP via the ICD. No part of the payload may extend aft of the payload interface plane without specific OSP approval. Incursions below the payload interface plane may be approved on a case-by-case basis after additional verification that the incursions do not cause any detrimental effects. Vertices for
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Figure 5-1. Standard 92 in. Fairing Envelope
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Orbital will provide all flight hardware and integration services necessary to attach nonseparating and separating payloads to Minotaur IV. Payload ground handling equipment is typically the responsibility of the payload contractor. All attachment hardware, whether Orbital or customer provided, must contain locking features consisting of locking nuts, inserts or fasteners. 5.2.1. Standard Non-Separating Mechanical Interface Orbital’s payload interface design provides a standard interface that will accommodate multiple payload configurations. Figure 5-2 illustrates the standard, non-separating payload mechanical interface. This is for payloads that provide their own separation system or payloads that will not separate. The interface is a standardized circular bolted interface common with the Evolved Expendable Launch Vehicle (EELV). The interface is a 62.01-inch diameter bolted interface. A butt joint with 121 holes (0.265-inch diameter) designed for ¼-inch fasteners is the payload mounting surface as shown in Figure 5-2. Alternate or multiple payload configurations can also be accommodated with the use of a bulkhead which allows flexibility in mounting patterns and configurations. 5.2.2. Orbital Supplied Mechanical Interface Control Drawing Orbital will provide a toleranced Mechanical Interface Control Drawing (MICD) to the payload contractor to allow accurate machining of the fastener holes. The Orbital provided MICD is the only approved documentation for drilling the payload interface. 5.3. Payload Electrical Interfaces The payload electrical interface supports battery charging, external power, discrete commands, discrete telemetry, analog telemetry, serial communication, payload separation indications, and up to 16 separate ordnance discretes. If an optional Orbital-provided
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separation system is utilized, Orbital will provide all the wiring through the separable interface plane. If the option is not exercised the customer will be responsible to provide the wiring from the spacecraft to the separation plane. 5.3.1. Payload Umbilical Interfaces The payload umbilical connector provides 60 wires from the ground to the spacecraft via a dedicated payload umbilical within the vehicle, as shown in Figure 5-3. The length of the internal umbilical is approximately 25 ft (7.62 m). The cabling from the LEV to the launch vehicle is approximately 130 ft (39.6 m). This umbilical is a dedicated pass through harness for ground processing support. It allows the payload command, control, monitor, and power to be easily configured per each individual user’s requirements. The umbilical wiring is configured as a one-to-one from the Payload/Minotaur IV interface through to the payload EGSE interface in the Launch Equipment Vault, the closest location for operating customer supplied payload EGSE equipment. It is a Launch Vehicle requirement that the payload provide two (2) separation loopback circuits on the payload side of the separation plane. These are typically wired into different separation connectors for redundancy. These breakwires are used for positive separation indication telemetry and initiation of the CCAM maneuver. Figure 5-4 details the pin outs for the standard interface umbilical. All wires are twisted, shielded pairs, and pass through the entire umbilical system, both vehicle and ground, as one-to-one to simplify and standardize the payload umbilical configuration requirements while providing maximum operational flexibility to the payload provider.
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Figure 5-2. Standard, Non-separating Payload Mechanical Interface
Figure 5-3. Payload Electrical Interface Block Diagram, With No Orbital Supplied Separation System
Figure 5-4. Payload 1:1 Umbilical Pin Outs
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5.3.2. Payload Interface Circuitry Standard interface circuitry passing through the payload-to-launch vehicle electrical connections are shown in Figure 5-5. This figure details the interface characteristics for launch vehicle commands, discrete and analog telemetry, separation loopbacks, pyro initiation, and serial communications interfaces with the launch vehicle avionics systems. 5.3.3. Payload Battery Charging Orbital provides the capability for remote controlled charging of payload batteries, using a customer provided battery charger. This power is routed through the payload umbilical cable. Up to 5.0 amperes per wire pair can be accommodated. The payload battery charger should be sized to withstand the line loss from the LEV to the spacecraft. 5.3.4. Payload Command and Control The Minotaur IV standard interface provides discrete sequencing commands generated by the launch vehicle’s Ordnance Driver Module (ODM) that are available to the payload as closed circuit opto-isolator command pulses of 5 A in lengths of 35 ms minimum. The total number of ODM discretes is sixteen (16) and can be used for any combination of (redundant) ordnance events and/or discrete commands depending on the payload requirements. 5.3.5. Pyrotechnic Initiation Signals Orbital provides the capability to directly initiate 16 separate pyrotechnic conductors through two dedicated MACH Ordnance Driver Modules (ODM). Each ODM provides for up to eight drivers capable of a 5 A, 100 ms, current limited pulse into a 1.5 ohm resistive load. All eight channels can be fired simultaneously with an accuracy of 1 ms between channels. In addition, the ODM channels can be utilized to trigger high impedance discrete events if required. Safing for all payload ordnance events will be accomplished through an Arm/Disarm (A/D) Switch.
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Figure 5-5. Minotaur IV Payload Electrical Interface Block Diagram 5.3.6. Payload Telemetry The baseline telemetry subsystem capability provides a number of dedicated payload discrete (bi-level) and analog telemetry monitors through dedicated channels in the vehicle encoder. Up to 24 channels will be provided with type and data rate being defined in the mission requirements document. In addition, a GCI610 will be utilized in the encoder stack for serial data ranging up to 600 Kbs if required. The GCI610 utilizes an IRIG standard RS422 driver interface for simplicity in payload interface definition. The payload serial and analog data will be embedded in the baseline vehicle telemetry format. For discrete monitors,
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the payload customer must provide the 5 Vdc source and the return path. The current at the payload interface must be less than 10 mA. Separation breakwire monitors can be specified if required. The number of analog channels available for payload telemetry monitoring is dependent on the frequency of the data. Payload telemetry requirements and signal characteristics will be specified in the Payload ICD and should not change once the final telemetry format is released at approximately L-6 months. Orbital will tape, archive, and reduce the data into a digital format for delivery to the payloaders for review. Due to the use of strategic assets, Minotaur IV telemetry is subject to the provisions of the Strategic Arms Reduction Treaty (START). START treaty provisions require that certain Minotaur IV telemetry be unencrypted and provided to the START treaty office for dissemination to the signatories of the treaty. The extent to which START applies to the payload telemetry will be determined by RSLP. Encrypted payload telemetry can be added as a non-standard service pending approval by RSLP and the START treaty office. 5.3.7. Payload Separation Monitor Loopbacks Separation breakwire monitors are required on both sides of the payload separation plane. With the Orbital provided separation systems, Minotaur IV provides three (3) separation loopbacks on the launch vehicle side of the separation plane for positive payload separation indication.
(programmable) and data transmission bit rates. The number of channels, sample rates, etc. will be defined in the Payload ICD. 5.3.9. Non-Standard Electrical Interfaces Non-standard services such as serial command and telemetry interfaces can be negotiated between OSP and the payload contractor on a mission-by-mission basis. The selection of the separation system could also impact the payload interface design and will be defined in the Payload ICD. 5.3.10. Electrical Launch Support Equipment Orbital will provide space for a rack of customer supplied EGSE in the LCR, or either of the on-pad equipment vaults. The equipment will interface with the launch vehicle/spacecraft through either the dedicated payload umbilical interface or directly through the payload access door. The payload customer is responsible for providing cabling from the EGSE location to the launch vehicle/spacecraft. Separate payload ground processing harnesses that mate directly with the payload can be accommodated through the payload access door(s) as defined in the Payload ICD. 5.4. Payload Design Constraints The following sections provide design constraints to ensure payload compatibility with the Minotaur IV system.
Minotaur IV also requires two (2) separate loopbacks on the payload side of the separation plane. These are used for telemetry indication of separation and also the initiation of the Stage 4 CCAM maneuver.
5.4.1. Payload Center of Mass Constraints Along the Y and Z-axes, the payload CG must be within 1.0 inch (3.8 cm) of the vehicle centerline. Payloads whose CG extend beyond the 1.0 inch lateral offset limit will require Orbital to verify the specific offsets that can be accommodated.
5.3.8. Telemetry Interfaces The standard Minotaur IV payload interface provides a 16Kbps RS-422/RS-485 serial interface for payload use with the flexibility to support a variety of channel/bit rate requirements, and provide signal conditioning, PCM formatting
5.4.2. Final Mass Properties Accuracy The final mass properties statement must specify payload weight to an accuracy of at least 1 lbm (0.5 kg), the center of gravity to an accuracy of at least 0.25 inch (6.4 mm) in each axis, and the
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products of inertia to an accuracy of at least 0.5 slug-ft2 (0.7 kg-m2) (see Figure 5-6). In addition, if the payload uses liquid propellant, the slosh frequency must be provided to an accuracy of 0.2 Hz, along with a summary of the method used to determine slosh frequency.
Figure 5-6. Payload Mass Properties Measurement Tolerance 5.4.3. Pre-Launch Electrical Constraints Prior to launch, all payload electrical interface circuits are constrained to ensure there is no current flow greater than 10 mA across the payload electrical interface plane. The primary support structure of the spacecraft shall be electrically conductive to establish a single point electrical ground. 5.4.4. Payload EMI/EMC Constraints The Minotaur IV avionics share the payload area inside the fairing such that radiated emissions compatibility is paramount. OSP places no firm radiated emissions limits on the payload other than the prohibition against RF transmissions within the payload fairing. Prior to launch, Orbital requires review of the payload radiated emission levels (MIL-STD-461, RE02) to verify overall launch vehicle EMI safety margin (emission) in accordance with MIL-E-6051. Payload RF transmissions are not permitted after fairing mate and prior to an ICD specified time after separation of the payload. An EMI/EMC analysis may be required to ensure RF compatibility. Payload RF transmission frequencies must be coordinated with Orbital and range officials to ensure non-interference with Minotaur IV and range transmissions. Additionally, the customer Release 1.1
must schedule all RF tests at the integration site with Orbital in order to obtain proper range clearances 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 25 Hz (based on Taurus). However, this value is effected significantly by other factors such as incorporation of a spacecraft isolation system and/or separation system. Therefore, the final determination of compatibility must be made on a mission-specific basis. 5.4.6. Payload Propellant Slosh A slosh model should be provided to Orbital 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 RSLP. 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 for Minotaur IV payloads. These are compliance documents and must be strictly followed. It is the responsibility of the customer to ensure that the payload meets all OSP, Orbital, and range imposed safety standards. Customers designing payloads that employ hazardous subsystems are advised to contact OSP early in the design process to verify compliance with system safety standards.
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6. MISSION INTEGRATION 6.1. Mission Management Approach The Minotaur IV program is managed through U.S. Air Force, Space and Missile Systems Center, Rocket Systems Launch Program (RSLP). RSLP serves as the primary point of contact for the payload customers for the Minotaur IV launch service. A typical integrated OSP organizational structure is shown in Figure 6-1. Open communication between RSLP , Orbital, and the customer, emphasizing timely transfer of data and prudent decision-making, ensures efficient launch vehicle/payload integration operations. 6.1.1. RSLP Mission Responsibilities The program office for all OSP missions is the RSLP . They are the primary Point of Contact (POC) for all contractual and technical coordination. RSLP contracts with Orbital to provide the Launch Vehicle and launch integration and separately with commercial Spaceports and/or Government Launch Ranges for launch site facilities and services. Once a mission is identified, RSLP will assign a Mission Manager to
coordinate all mission planning and contracting activities. RSLP is supported by Northrop Grumman and other associate contractors for technical and logistical support, particularly utilizing their extensive expertise and background knowledge of the Peacekeeper booster and subsystems. 6.1.2. Orbital Mission Responsibilities As the launch vehicle provider, Orbital’s responsibilities fall into four primary areas: a. Launch Vehicle Program Management b. Mission Management c. Engineering d. Launch Site Operations Orbital assigns a Mission Manager to manage the launch vehicle technical and programmatic interfaces for a particular mission. The Orbital Mission Manager is the single POC for all aspects of a specific mission. This person has overall program authority and responsibility to ensure that payload requirements are met and that the appropriate launch vehicle services are provided. The Orbital Mission Manager will jointly chair the
Figure 6-1. Typical Integrated OSP Organizational Structure Release 1.1
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Mission Integration Working Groups (MIWGs) with the RSLP Mission Manager. The Mission Managers responsibilities include detailed mission planning, payload integration services, systems engineering, mission-peculiar design and analyses coordination, payload interface definition, launch range coordination, integrated scheduling, launch site processing, and flight operations. 6.2. Mission Planning and Development OSP will assist the customer with mission planning and development associated with Minotaur IV launch vehicle systems. These services include interface design and configuration control, development of integration processes, launch vehicle analyses, facilities planning, launch campaign planning to include range services and special operations, and integrated schedules. The procurement, analysis, integration and test activities required to place a customer’s payload into orbit are typically conducted over a 20 month standard sequence of events called the Mission Cycle. This cycle normally begins 18 months before launch, and extends to 8 weeks after launch. Once contract authority to proceed is received, the Mission Cycle is initiated. The contract option designates the payload, launch date, and basic mission parameters. In response, the Minotaur IV Program Manager designates an Orbital Mission Manager who ensures that the launch service is supplied efficiently, reliably, and on-schedule. The typical Mission Cycle interweaves the following activities: a. Mission management, document exchanges, meetings, and formal reviews required to coordinate and manage the launch service. b. Mission analyses and payload integration, document exchanges, and meetings. c. Design, review, procurement, testing and integration of all mission-peculiar hardware and software.
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d. Range interface, safety, and flight operations activities, document exchanges, meetings and reviews. Figure 6-2 details the typical Mission Cycle for a specific launch and how this cycle folds into the Orbital vehicle production schedule with typical payload activities and milestones. A typical Mission Cycle is based on an 18 month interval between mission authorization and launch. This interval reflects the OSP contractual schedule and has been shown to be an efficient schedule based on Orbital’s Minotaur, Taurus and Pegasus program experience. However, OSP is flexible to negotiate either accelerated cycles, which take advantage of the Minotaur IV/Pegasus/Minotaur/ Taurus multi-customer production sets, or extended cycles required by unusual payload requirements, such as extensive analysis or complex payload-launch vehicle integrated designs or tests or funding limitations. 6.3. Mission Integration Process 6.3.1. Integration Meetings The core of the mission integration process consists of a series of Mission Integration and Range Working Groups (MIWG and RWG, respectively). The MIWG has responsibility for all physical interfaces between the payload and the launch vehicle. As such, the MIWG creates and implements the Payload-to-Minotaur IV ICD in addition to all mission-unique analyses, hardware, software, and integrated procedures. The RWG is responsible for the areas of launch site operations; range interfaces; safety review and approval; and flight design, trajectory, and guidance. Documentation produced by the RWG includes all required range and safety submittals. Working Group membership consists of the Mission Manager and representatives from Minotaur IV engineering and operations organizations, as well as their counterparts from the customer organization. While the number of meetings, both formal and informal, required to develop and implement the mission integration
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Figure 6-2. Typical Mission Integration Schedule process will vary with the complexity of the spacecraft, quarterly meetings are typical. 6.3.2. Mission Design Reviews (MDR) Two mission-specific design reviews will be held to determine the status and adequacy of the launch vehicle mission preparations. They are designated MDR-1 and MDR-2 and are typically held 6 months and 13 months, respectively, after authority to proceed. They are each analogous to Preliminary Design Reviews (PDRs) and Critical Design Reviews (CDRs), but focus primarily on mission-specific elements of the launch vehicle effort. 6.3.3. Readiness Reviews During the integration process, reviews are held to provide the coordination of mission participants and management outside of the regular contact of the Working Groups. Due to the
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variability in complexity of different payloads and missions, the content and number of these reviews can be tailored to customer requirements. As a baseline, Orbital will conduct two readiness reviews as described below. a. Mission Readiness Review — Conducted within 1 month of launch, the Mission Readiness Review (MRR) provides a pre-launch assessment of integrated launch vehicle/payload/facility readiness prior to committing significant resources to the launch campaign. b. Launch Readiness Review — The Launch Readiness Review (LRR) is conducted at L-1 day and serves as the final assessment of mission readiness prior to activation of range resources on the day of launch.
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6.4. Documentation Integration of the payload requires detailed, complete, and timely preparation and submittal of interface documentation. As the launch service provider, RSLP is the primary communication path with support agencies, which include—but are not limited to—the various Range support agencies and U.S. Government agencies such as the U.S. Department of Transportation and U.S. State Department. Customer-provided documents represent the formal communication of requirements, safety data, system descriptions, and mission operations planning. The major products and submittal times associated with these organizations are divided into two areas— those products that are provided by the customer, and those produced by Orbital. 6.4.1. Customer-Provided Documentation Documentation produced by the customer is detailed in the following paragraphs. 6.4.1.1. Payload Questionnaire The Payload Questionnaire is designed to provide the initial definition of payload requirements, interface details, launch site facilities, and preliminary safety data to OSP. The customer shall provide a response to the Payload Questionnaire form (Appendix A), or provide the same information in a different format, in time to support the Mission Kickoff Meeting. The customer’s responses to the payload questionnaire define the most current payload requirements and interfaces and are instrumental in Orbital’s preparation of numerous documents including the ICD, Preliminary Mission Analysis, and launch range documentation. Additional pertinent information, as well as preliminary payload drawings, should also be included with the response. Orbital understands that a definitive response to some questions may not be feasible. These items are defined during the normal mission integration process. 6.4.1.2. Payload Mass Properties Payload mass properties must be provided in a timely manner in order to support efficient launch vehicle trajectory development and dynamic Release 1.1
analyses. Preliminary mass properties should be submitted as part of the MRD at launch vehicle authority to proceed. Updated mass properties shall be provided at predefined intervals identified during the initial mission integration process. Typical timing of these deliveries is included in Figure 6-2. 6.4.1.3. Payload Finite Element Model A payload mathematical model is required for use in Orbital’s preliminary coupled loads analyses. Acceptable forms include either a Craig-Bampton model valid to 120 Hz or a NASTRAN finite element model. For the final coupled loads analysis, a test verified mathematical model is desired. 6.4.1.4. Payload Thermal Model for Integrated Thermal Analysis An integrated thermal analysis can be performed for any payload as a non-standard service. A payload thermal model will be required from the payload organization for use in Orbital’s integrated thermal analysis if it is required. The analysis is conducted for three mission phases: a. Prelaunch ground operations; b. Ascent from lift-off until fairing jettison; and c. Fairing jettison through payload deployment. Models must be provided in SINDA format. There is no limit on model size although turnaround time may be increased for large models. 6.4.1.5. Payload Drawings Orbital prefers electronic versions of payload configuration drawings to be used in the mission specific interface control drawing, if possible. Orbital will work with the customer to define the content and desired format for the drawings. 6.4.1.6. Program Requirements Document (PRD) Mission Specific Annex Inputs To obtain range support, a PRD must be prepared. This document describes requirements needed to generally support the Minotaur IV launch vehicle. For each launch, an annex is submitted to specify the range support needed to
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meet the mission’s requirements. This annex includes all payload requirements as well as any additional Minotaur IV requirements that may arise to support a particular mission. The customer completes all appropriate PRD forms for submittal to Orbital. 6.4.1.6.1. Launch Operations Requirements (OR) Inputs To obtain range support for the launch operation and associated rehearsals, an OR must be prepared. The customer must provide all payload pre-launch and launch day requirements for incorporation into the mission OR. 6.5. Safety 6.5.1. System Safety Requirements In the initial phases of the mission integration effort, regulations and instructions that apply to spacecraft design and processing are reviewed. Not all safety regulations will apply to a particular mission integration activity. Tailoring the range requirements to the mission unique activities will be the first step in establishing the safety plan. OSP has three distinctly different mission approaches affecting the establishment of the safety requirements: a. Baseline mission: Payload integration and launch operations are conducted at VAFB, CA b. Campaign/VAFB Payload Integration mission: Payload integration is conducted at VAFB and launch operations are conducted from a non-VAFB launch location. c. Campaign/Non-VAFB Payload Integration mission: Payload integration and launch operations are conducted at a site other than VAFB. For the baseline and VAFB Payload Integration missions, spacecraft prelaunch operations are conducted at Government Furnished Property (GFP) Payload Processing Facility (PPF). For campaign style missions, the spacecraft prelaunch operations are performed at the desired launch site. Release 1.1
Before a spacecraft arrives at the processing 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 Requirements for baseline and VAFB Payload Integration 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. Processing and launch site ranges discourage the use of waivers and variances. Furthermore, approval of such waivers cannot be guaranteed. 6.5.2. System Safety Documentation For each Minotaur IV 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 respective systems should be included. Major categories of hazardous systems are ordnance devices, radioactive materials, propellants, pressurized systems, toxic materials, cryogenics, and RF radiation. Procedures relating to these systems as well as any procedures relating to lifting operations or battery operations should be prepared for safety review submittal. OSP will provide this information to the appropriate safety offices for approval.
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Section 7.0 – Ground and Launch Operations
7. GROUND AND LAUNCH OPERATIONS 7.1. Minotaur IV/Payload Integration Overview The processing of the Minotaur IV utilizes many of the same proven techniques developed for the Pegasus, Taurus and Minotaur launch vehicles. This minimizes the handling complexity for both vehicle and payload. 7.2. Ground And Launch Operations Ground and launch operations are conducted in three major phases: a. Launch Vehicle Integration — Assembly and test of the Minotaur IV vehicle b. Payload Processing/Integration — Receipt and checkout of the satellite payload, followed by integration with Minotaur IV fairing and Payload Adapter Module (PAM) and verification of interfaces c. Launch Operations — Includes transport to the launch pad, final integration, checkout, arming and launch. 7.2.1. Launch Vehicle Integration Orbital will process all Minotaur IV vehicles according to a flow similar to that implemented for the Minotaur and Taurus vehicles. All launch vehicle motors, parts and completed subassemblies are delivered to the launch Vehicle Processing Facility (VPF) from Orbital’s Chandler production facility, the assembly/motor vendor or the Government. Figure 7-1 depicts the typical flow of hardware from the factory to the launch site. Flowcharts of the field processing are shown in Figure 7-2. 7.2.1.1. Planning and Documentation Minotaur IV integration and test activities are controlled by a comprehensive set of Work Packages (WPs) that describe and document every aspect of integrating and testing Minotaur IV and its payload. All testing and integration activities are scheduled by work package number on a daily activity schedule updated and distributed daily during field operations. This schedule is maintained by Orbital and serves as Release 1.1
the master document communicating all activities planned at the field site. The schedule contains notations regarding the status of the work package document and hardware required to begin the operation. Mission-specific work packages are created for mission-unique or payload-specific procedures. Any discrepancies encountered are recorded on a Discrepancy Report and dispositioned as required. All activities are in accordance with Orbital’s ISO 9001 certification. 7.2.1.2. GCA/Orion 38 Integration and Test Activities The GCA will undergo subsystem level testing at Orbital’s Chandler facility prior to being shipped to the field site. The GCA and the Stage 4 Orion 38 motor are then delivered to the launch vehicle processing facility located at VAFB. Upon arrival at VAFB these components/sub-assemblies will undergo a thorough inspection and subsystem level checkout. At this time range certification of Range Tracking System (RTS) and Flight Termination System (FTS) devices will be performed. The components will be reinstalled and in-vehicle testing of the RTS and FTS systems will be performed. After the completion of subsystem level testing the Orion 38 motor is integrated into the GCA to form the Stage 4 assembly. 7.2.1.3. PK Motor Integration and Test Activities The PK Stage 1, 2 and 3 motors are delivered to the launch vehicle processing facility where they undergo checkout and testing. Once integration is complete, a booster confidence test will be conducted. 7.2.1.4. Mission Simulation Tests Orbital will run three Mission Simulator Tests (MST) to verify the functionality of launch vehicle hardware, and software. The Mission Simulation Tests use the actual flight software and simulate a “fly to orbit” scenario using simulated Inertial Navigation System (INS) data. This will allow the test to proceed throughout all mission phases
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Figure 7-1. Hardware Flow – Factory to Launch Site
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Figure 7-2. SLV Processing Flow recording vehicle performance data. The data will be compared to simulations performed in the factory software laboratory using an identical copy of the flight software. Orbital will use GFP PK nozzle assembly simulators to perform all mission simulations. These components will provide a realistic assessment of booster performance during the testing operations. After a thorough data review of all telemetry parameters, the test configuration is disassembled and setup for payload integration begins. 7.2.1.5. Booster Assembly Stacking/Launch Pad Preparation After completion of the MST, the booster assembly (Stages 1, 2 and 3) and the stage 4 assembly (Orion 38 integrated with the GCA) are transported to the launch facility. Figure 7-3 shows a pictorial representation of the processing flow. Prior to the arrival of the PK boosters, the site is prepared for launch operations with the Release 1.1
installation of the launch stool. This stool is the same design used for Orbital’s Taurus SLV. It supports a flat pad launch of a full PK booster assembly with front section and fourth stage motor options. After stool installation, the fixed scaffolding installation is performed. This scaffolding provides access to the base ring of the PK Stage 1 motor during integration activities. Once the booster arrives at the launch site, it is then lifted and emplaced onto the launch stool. Each motor assembly will be individually stacked using a process developed for handling Taurus Stage 0 motors. Scaffolding integration continues as the booster stages are mated. The Stage 4 assembly is shipped in the vertical configuration to the launch facility for payload integration. 7.2.2. Payload Processing/Integration Payloads normally undergo initial checkout and preparation for launch at an Air Force payload
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Figure 7-3. Minotaur IV Processing Flow processing facility (PPF) or commercial facilities at VAFB. After arrival at the PPF, the payload completes its own independent verification and checkout prior to beginning integrated processing with Minotaur IV fairing and Payload Adapter Module (PAM). The Minotaur IV fairing and PAM will be delivered to the payload processing facility for encapsulation of the payload. The fairing and PAM provide a sealed enclosure which protects the payload and provides a structure to facilitate transportation to the launch facility. After enclosure of the payload in the fairing, the assembly is shipped in the vertical configuration to the launch facility for a pre-installation verification test. 7.2.2.1. Payload Propellant Loading Payloads utilizing integral propulsion systems with propellants such as hydrazine can be loaded and secured through coordinated Orbital and contractor arrangements for use of the propellant Release 1.1
loading facilities in the VAB. standard service.
This is a non-
7.2.2.2. Final Vehicle Integration and Test After successful completion of payload mate/fairing closeout the completed front section assembly (Minotaur IV Stage 4 assembly integrated with the payload assembly) will then be lifted in vertical configuration atop the booster assembly. Figure 7-4 illustrates the vertical lifting operation performed on a Taurus front section. Final post mates checks of the booster assembly and front section assembly interface are then conducted. A final systems verification test, similar to the previous MST, is then performed. At this point the vehicle is ready for final Range interface tests and launch readiness.
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Figure 7-4. Minotaur IV Upper Stack Assembly will be Vertically Integrated to Minotaur IV Booster Assembly in a Similar Manner to Taurus Upper Stack
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7.3.1. Launch Control Organization The Launch Control Organization is split into two groups: the Management group and the Technical group. The Management group consists of senior range personnel and Mission Directors/Managers for the launch vehicle and payload. The Technical Group consists of the personnel responsible for the execution of the launch operation and data review/assessment for the Payload, the Launch Vehicle and the Range. The Payload’s members of the technical group are engineers who provide technical representation in the control center. The Launch Vehicle’s members of the technical group are engineers who prepare the Minotaur IV for flight, review and assess data that is displayed in the Launch Control Room (LCR) and provide technical representation in the LCR and in the Launch Operations Control Center (LOCC). The Range’s members of the technical group are personnel that maintain 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.
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8. OPTIONAL ENHANCED CAPABILITIES The OSP launch service is structured to provide a baseline vehicle configuration which is then augmented with optional enhancements to meet the unique needs of individual payloads. The baseline vehicle capabilities are defined in the previous sections and the optional enhanced capabilities are defined below. The enhanced options allow customization of launch support and accommodations to the Minotaur IV designs on an efficient, “as needed” basis. 8.1. Mechanical Interface and Separation System Enhancements 8.1.1. Separation Systems Various separation systems can be provided or accommodated to meet mission-unique requirements. As a baseline option, Orbital offers an optional payload separation system that is flight proven on Taurus. The separation system is manufactured for Orbital by SAAB Ericson Space (SES). SES has extensive experience in supplying separation systems for a wide range of launch vehicles and payloads. This system is based on a design that has flown over 30 times with 100% success. The baseline separation system, shown in Figure 8-1, has a standardized 38.81 inch bolt pattern. It is a marmon clamp design employing two aluminum interface rings that are clamped by dual, semi-circular stainless steel clamp bands with aluminum clamp shoes. Each of the two retention bolts is severed by a redundantly initiated bolt cutter. The separation ring to which the payload attaches is supplied with through holes and the separation system is mated to the spacecraft during processing at the PPF. The weight of hardware separated with the payload is approximately 8.7 lbm (13.95 kg). Orbitalprovided attachment bolts to this interface can be inserted from either the launch vehicle or the payload side of the interface (NAS630xU, dash number based on payload flange thickness). The weight of the bolts, nuts, and washers connecting
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the separation system to the payload is allocated to the separation system and included in the launch vehicle mass. Separation velocity is provided by up to eight matched spring actuators. The spring assemblies may be tailored to mitigate the effects of payload CG offset, controlling tip-off within 5 deg/sec. Tipoff rates are highly dependent on payload mass properties, but are typically on the order 1 deg/sec. Preliminary and final mission-specific tipoff analyses are conducted for each payload using Orbital’s computer simulation dynamic analysis tools. If non-standard separation velocities are needed, different springs may be substituted on a mission-specific basis as a non-standard service. Other separation systems can supplied on a mission-specific basis.
also
be
8.1.2. Additional Fairing Access Doors Additional access doors can be provided to accommodate unique payload requirements. The standard door size is 18 in. by 24 in. (45.7 cm by 61.0 cm). Access doors of non-standard size can also be provided as necessary. Orbital performs structural analyses to verify the acceptability of the mission-specific door configuration. 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. The 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 and six Taurus missions. The typical Minotaur configuration is shown in Figure 8-2. 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 overall payload center of gravity loads and component or subsystem
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Figure 8-1. 38-in. Separation System Option responses. Typically the system will reduce transient loads to approximately 50% of the level they would be without the system. The exact results can be expected to vary for each particular spacecraft and with location on the spacecraft. Generally, a beneficial reduction in shock and vibration will also be provided. The isolation system does impact overall vehicle performance (by approximately 20 to 40 lb [9 to 18 kg]) and the available payload dynamic envelope by up to 4 inches (10.16 cm) axially and up to 1.0 inch (2.54 cm) laterally.
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8.2. Performance Enhancements 8.2.1. Insertion Accuracy Enhanced insertion accuracy or support for multiple payload insertion can be provided as an enhanced option utilizing the Hydrazine Auxiliary Propulsion System (HAPS). The common usage of the Orion 38 makes the flight proven HAPS design directly applicable to the Minotaur IV usage, as shown in Figure 8-3. Orbital insertion accuracy can typically be improved to ±10 nmi (±18.5 km) or better in each apse and ±0.05 deg in inclination. For orbits above 324 nmi (600 km), the
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Figure 8-2. Soft Ride Payload Isolation System as Integrated on Minotaur LV
HAPS also permits injection of shared payloads into different orbits. HAPS, which is mounted inside the Avionics Structure, consists of a hydrazine propulsion subsystem and a Stage 4 separation subsystem. After burnout and separation from the Stage 4 motor, the HAPS hydrazine thrusters provide additional velocity and both improved performance and precise orbit injection. The HAPS propulsion subsystem consists of a centrally mounted tank containing approximately 130 lbm (59 kg) of hydrazine, helium pressurization gas, and three fixed, axially pointed thrusters. The hydrazine tank contains an integral bladder which will support multiple restarts.
HAPS can also increase payload mass by approximately 50 to 250 lbm (22.7 to 113 kg), depending on the orbit. Specific performance capability associated with the HAPS can be provided by contacting the OSP program office.
8.2.2. Star 48 Stage 4 The modular design of Orbital’s GCA and integrating structures provides great flexibility in accommodating alternative Stage 4 propulsion systems. As one low risk example, a optional
Figure 8-3. Hydrazine Auxiliary Propulsion System (HAPS) Used to Provide Insertion Accuracy
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configuration using the more powerful ATK Thiokol Star-48 motor is shown in Figure 8-4. The only modifications required to accommodate this change are a modified Motor Adapter Cone (MAC) with the Star 48 forward interface and a longer 3/4 interstage to allow room for the increased motor length. Other alternative motors can be similarly adopted.
100,000 or Class 10,000). Nitrogen purge is used in conjunction with the Conditioned Air option to provide mission-specific localized cooling and/or dry nitrogen environments to satisfy unique payload environmental requirements. 8.3.2. Nitrogen Purge Continuous clean dry nitrogen inside the shroud during vehicle processing from payload encapsulation to launch is available as an option. Dry clean nitrogen purge can be provided to the payload at a Class 10,000 environment for continuous purge of the payload after fairing encapsulation until lift-off. The capability was demonstrated on the Minotaur MightySat mission with the exception of purge during transportation. A nitrogen cooling system is already provided on every SLV mission to spot-cool sensitive electronic boxes. Flow adjustments for cooling versus purge would be changed back and forth to accommodate both. If nitrogen purge is required during transport, the only additional items needed would be a minor addition of a nitrogen bottle, regulator, and mounting hardware.
Figure 8-4. Orion 38 Stage 4 Motor can be Replaced with a Star-48 to Provide Increased Performance 8.3. Environmental Control Options 8.3.1. Conditioned Air Conditioned air can be provided within the fairing volume using an Environmental Control System (ECS) via a “fly-out” duct that is retracted at launch. Temperature and humidity is regulated within the limits specified in the Payload ICD. A filter is installed to provide a Class 100,000 environment, typically. The Nitrogen Purge (Section 4.6.3) and Enhanced Contamination Control (Section 4.7 and 8.3.3) enhancements complement this capability. Upon exercise of the Enhanced Cleanliness option, a certified HEPA filter is used in the input duct to assure the necessary low particulate environment (Class Release 1.1
The system distribution lines are routed across the payload interface plane and/or along the inner surface of the shroud or fairing. If required for spot cooling of a payload component, Orbital will provide GN2 flow to localized regions in the fairing. The GN2 will meet Grade B specifications, as defined in MIL-P-27401C and can be regulated to at least 5 scfm. The system’s regulators are set to a desired flow rate during pre-launch processing. Payload purge requirements are controlled and documented via the launch vehicle to payload ICD. Payload purge requirements must be coordinated with Orbital via the ICD to ensure that the requirements can be achieved. 8.3.3. Enhanced Contamination Control Understanding that some payloads have requirements for enhanced cleanliness, OSP offers a contamination control option, which is
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composed of the elements in the following sections (which is also discussed in Section 4.7). Minotaur IV customers can also coordinate combinations of the elements listed below to meet the unique needs of their payloads. 8.3.3.1. High Cleanliness Integration Environment (Class 10K or 100K) With enhanced contamination control, a soft walled clean room can be provided to ensure a FED-STD-209 Class M6.5 (100,000) or Class M5.5 (10,000) environment during all payload processing activities up to fairing encapsulation. The soft walled clean room and anteroom(s) utilize HEPA filter units to filter the air and hydrocarbon content is maintained at 15 ppm or less. The payload organization is responsible for providing the necessary clean room garments for payload staff as well as vehicle staff that need to work inside the clean room. 8.3.3.2. Fairing Surface Cleanliness Options The inner surface of the fairing and payload cone assemblies can be cleaned to cleanliness criteria which ensures no particulate matter visible with normal vision when inspected from 6 to 18 inches under 100 ft-candle incident light. The same will be true when the surface is illuminated using black light, 3200 to 3800 Angstroms (Visibly Clean Plus Ultraviolet). In addition, Orbital can ensure that all materials used within the encapsulated volume have outgassing characteristics of less than 1.0% TML and less than 0.1% CVCM. Items that do not meet these levels can be masked to ensure they are encapsulated and will have no significant effect on the payload. 8.3.3.3. High Cleanliness Fairing Environment With the enhanced contamination control option, Orbital provides an ECS from payload encapsulation until just prior to vehicle lift-off. The ECS continuously purges the fairing volume with clean filtered air. Orbital’s ECS incorporates a HEPA filter unit to provide FED-STD-209 Class M5.5 (10,000) air. Orbital monitors the supply air for particulate matter via a probe installed Release 1.1
upstream of the fairing inlet duct prior to connecting the air source to the payload fairing. 8.3.4. Launch Pad Environmental Control For launch sites without gantries for environmental control or vehicle access, optional Quick Erect Scaffold® (QES) will protect the vehicle from the environments, maintain temperatures within 60 to 100 °F (in conjunction with a thermal blanket) and provide access to the vehicle for launch pad operations (see Figure 8-5). As the name implies, QES can be rapidly assembled and is highly adaptable for accommodating different vehicle configurations. This scaffolding has been previously demonstrated on Minotaur. 8.3.4.1. Booster Temperature Control The thermal blanket design successfully used during Minotaur missions can be used to maintain the PK booster operating temperature within the limits of 60 to 100 °F. The thermal blanket is
Figure 8-5. Mobile Scaffolding for Environmental Control Demonstrated on Minotaur Missions
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constructed of outer PVC material with an inner insulating liner. It is a four piece cover with Velcro seams running along the length of the boosters. Integral inflatable manifold tubes space the blanket away from the booster and provide space for ducting conditioned air for the boosters. The baseline blanket design only covers the Stage 1, 2 and 3 boosters. 8.4. Enhanced Telemetry Options OSP can provide mission specific instrumentation and telemetry components to support additional payload or experiment data acquisition requirements. Telemetry options include additional payload-dedicated bandwidth and GPS-based precision navigation data. 8.4.1. Enhanced Telemetry Bandwidth Enhanced mission specific instrumentation and telemetry can be provided, supplying a dedicated telemetry link to support additional payload or experiment data acquisition requirements. A baseline data rate of 1 Mbps is available, however, maximum data rates depend on the mission coverage required and the launch range receiver characteristics and configuration. The enhanced telemetry option was demonstrated on both inaugural Minotaur TLV and SLV missions. 8.4.2. Enhanced Telemetry Instrumentation To support the higher data rate capability in Section 8.4.1, enhanced telemetry instrumentation can be provided. The instrumentation can include strain gauges, temperature sensors, accelerometers, analog data, and digital data configured to mission-specific requirements. This capability was successfully demonstrated on the inaugural OSP-SLV mission. 8.4.3. Navigation Data Precision navigation data using an independent Global Positioning System (GPS) receiver and telemetry link is available as an enhanced option. This option utilizes Orbital’s flight proven GPS Position Beacon (GPB) to provide missile state data for range safety and
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provides better than 100 m position accuracy with 10 Hz data rate. This capability was successfully demonstrated on the inaugural Minotaur mission. 8.5. Shared Launch Accommodations Minotaur IV is uniquely capable of providing launches of multiple satellite payloads, leveraging RSLP and Orbital’s extensive experience in integrating and launching multiple payloads. Multiple spacecraft configurations have been flown on many of Orbital’s Pegasus, Taurus and Minotaur missions to date. A number of different structural configurations have been developed for dual payloads, one is shown in Figure 8-6. Because of the modular nature of the structures, dual payload configurations can be easily accommodated by the Minotaur IV structural design.
Figure 8-6. Modular Minotaur IV Structural Design Easily Accommodates Multiple Payloads
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Appendix A
APPENDIX A PAYLOAD QUESTIONNAIRE
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SATELLITE IDENTIFICATION FULL NAME: ACRONYM: OWNER/OPERATOR: INTEGRATOR(s): ORBIT INSERTION REQUIREMENTS* SPHEROID
Standard (WGS-84, Re = 6378.137 km) Other:
ALTITUDE Insertion Apse: nmi ___ or... Semi-Major Axis:
Opposite Apse: ±
__
km
nmi km
± Eccentricity:
nmi ___ INCLINATION
±
__
≤e≤
km
± deg
ORIENTATION Argument of Perigee:
Longitude of Ascending Node (LAN): ±
deg
±
deg
Right Ascension of Ascending Node (RAAN): ±
deg
...for Launch Date: * Note: Mean orbital elements
LAUNCH WINDOW REQUIREMENTS NOMINAL LAUNCH DATE: OTHER CONSTRAINTS (if not already implicit from LAN or RAAN requirements, e.g., solar beta angle, eclipse time constraints, early on-orbit ops, etc):
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Appendix A
GROUND SUPPORT EQUIPMENT Describe any additional control facilities (other than the baseline Support Equipment Building (SEB) and Launch Equipment Vault (LEV)) which the satellite intends to use:
SEB
Describe (in the table below) Satellite EGSE to be located in the LSV. [Note: Space limitations exist in the SEB, 350 ft umbilical cable length to spacecraft typical] Equipment Name / Type
Approximate Size (LxWxH)
Power Requirements
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................................
..................................................................
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................................
..................................................................
...............................................
................................
..................................................................
...............................................
................................
..................................................................
...............................................
................................
Is UPS required for equipment in the SEB? Is Phone/Fax connection required in the SEB? LEV
Yes / No Yes / No
Circle: Phone / FAX
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] Approximate Size (LxWxH) Power Requirements Equipment Name / Type ..................................................................
...............................................
................................
..................................................................
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Is UPS required for equipment in the LEV? Is Ethernet connection between SEB and LEV required?
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EARLY ON-ORBIT OPERATIONS Briefly describe the satellite early on-orbit operations, e.g., event triggers (separation sense, sun acquisition, etc), array deployment(s), spin ups/downs, etc:
SATELLITE SEPARATION REQUIREMENTS ACCELERATION Longitudinal:
=
g’s
Lateral:
=
g’s
VELOCITY Relative Separation Velocity Constraints: ANGULAR RATES Longitudinal: (pre-separation) ±
deg/sec
ANGULAR RATES Longitudinal: (post-separation) ±
deg/sec
Pitch:
±
deg/sec
Yaw:
±
deg/sec
Pitch:
±
deg/sec
Yaw:
±
deg/sec
ATTITUDE Describe Pointing Requirements Including Tolerances: (at deployment)
SPIN UP Longitudinal Spin Rate: OTHER
±
deg/sec
Describe Any Other Separation Requirements:
SPACECRAFT COORDINATE SYSTEM Describe the Origin and Orientation of the spacecraft reference coordinate system, including its orientation with respect to the launch vehicle (provide illustration if available):
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SPACECRAFT PHYSICAL DIMENSIONS STOWED CONFIGURATION
Length/Height:
Diameter: in
cm
in
cm
Other Pertinent Dimension(s):
Describe any appendages/antennas/etc which extend beyond the basic satellite envelope:
ON-ORBIT Describe size and shape: CONFIGURATION
If available, provide dimensioned drawings for both stowed and on-orbit configurations.
SPACECRAFT MASS PROPERTIES* PRE-SEPARATION Mass: Xcg: Ycg: Zcg: POST-SEPARATION Mass: (non-separating adapter remaining with launch vehicle) Xcg: Ycg: Zcg:
Inertia units: lbm kg Ixx: in cm Izz: in cm Iyz: in cm Inertia units: lbm kg Ixx: in cm Izz: in cm Iyz: in cm
lbm-in2
kg-m2
Iyy: Ixy: Ixz: lbm-in2
kg-m2
Iyy: Ixy: Ixz:
* Stowed configuration, spacecraft coordinate frame
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ASCENT TRAJECTORY REQUIREMENTS Free Molecular Heating at Fairing Separation: (Standard Service: = 360 Btu/ft2/hr)
Btu/ft2/hr FMH =
W/m2
Fairing Internal Wall Temperature (Standard Service: = 200°F)
deg F T =
deg C
Dynamic Pressure at Fairing Separation: (Standard Service: = 0.01 lbf /ft2)
lbf /ft2 q =
N/m2
Ambient Pressure at Fairing Separation: (Standard Service: = 0.3 psia)
lbf /in2 P =
N/m2
Maximum Pressure Decay During Ascent: (Standard Service: = 0.6 psia)
lbf /in2/sec ΔP =
N/m2/sec
Thermal Maneuvers During Coast Periods: (Standard Service: none) SPACECRAFT ENVIRONMENTS THERMAL Spacecraft Thermal Dissipation, Pre-Launch Encapsulated: DISSIPATION Approximate Location of Heat Source: TEMPERATURE Temperature Limits During Ground/Launch Operations:
Watts
Max Min (Standard Service is 55°F to 80°F)
deg F deg F
deg C deg C
Component(s) Driving Temperature Constraint: Approximate Location(s): HUMIDITY Relative Humidity: Max % % Min
Dew Point: Max deg F Min deg F (Standard Service is 37 deg F)
or,
deg C deg C
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:
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Lateral =
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ELECTRICAL INTERFACE Bonding Requirements: Are Launch Vehicle Supplied Pyro Commands Required?
Yes / No If Yes, magnitude: (Standard Service is 10 amps for 100 msec)
Are Launch Vehicle Supplied Discrete Commands Required? Yes / No
amps for
msec
If Yes, describe:
Is Electrical Access to the Satellite Required... After Encapsulation? at Launch Site Yes / No
Yes / No
Is Satellite Battery Charging Required... After Encapsulation? at Launch Site? Yes / No
Yes / No
Is a Telemetry Interface with the Launch Vehicle Flight Computer Required?
Yes / No
If Yes, describe:
Other Electrical Requirements:
Please complete attached sheet of required pass-through signals. RF RADIATION Time After Separation Until RF Devices Are Activated: (Note: Typically, no spacecraft radiation is allowed from encapsulation until 30 minutes after liftoff.) Frequency:
MHz
Power:
Watts
Location(s) on Satellite (spacecraft coordinate frame): Longitudinal
in
cm
Clocking (deg), Describe:
Longitudinal
in
cm
Clocking (deg), Describe:
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REQUIRED PASS-THROUGH SIGNALS Item #
Pin
Signal Name
From LEV
To Satellite
Shielding
Max Current (amps)
Total Line Resistance (ohms)
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28
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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
Flatness Requirements for Sep System or Mating Surface of Launch Vehicle:
FAIRING Payload Fairing Access Doors (spacecraft coordinate frame): ACCESS Longitudinal in cm Clocking (deg), Describe: Longitudinal
in
cm
Clocking (deg), Describe:
Note: Standard Service is one door DYNAMICS Spacecraft Natural Frequency: Axial Recommended: OTHER
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Hz
Lateral
> TBD Hz
Hz > TBD Hz
Other Mechanical Interface Requirements:
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Copyright 2006 by Orbital Sciences Corporation. All Rights Reserved.
Minotaur IV User’s Guide
Revision Summary
REVISION SUMMARY VERSION DOCUMENT
DATE
CHANGE
PAGE
1.0
TM-17589
Jan 2005
Initial Release
All
1.1
TM-17589A
Jan 2006
General nomenclature, history, and administrative updates (no technical updates)
All
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•
Updated launch history
•
Corrected contact information
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Preface
This User’s Guide is intended to familiarize payload mission planners with the capabilities of the Orbital Suborbital Program ll (OSP-2) Minotaur IV Space Launch Vehicle (SLV) launch service. The information provided in this user’s guide is for initial planning purposes only. Information for development/design is determined through mission specific engineering analyses. The results of these analyses are documented in a mission-specific Interface Control Document (ICD) for the payloader organization to use in their development/design process. This document provides an overview of the Minotaur IV system design and a description of the services provided to our customers.
USAF SMC Det 12/RP 3548 Aberdeen Ave SE Kirtland AFB, NM 87117-5778 Telephone: Fax:
(505) 846-8957 (505) 846-5152
Additional copies of this User's Guide and Technical information may also be requested from Orbital at: Orbital Suborbital Program - Mission Development Orbital Sciences Corporation Launch Systems Group 3380 S. Price Road Chandler, AZ 85248 Telephone: E-mail:
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Preface
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Table of Contents
PAGE 1.
INTRODUCTION................................................................................................................................. 1
2.
MINOTAUR IV LAUNCH SERVICE .................................................................................................... 3 2.1.
Minotaur IV Launch System Overview.......................................................................................... 3
2.2.
Minotaur IV Launch Service.......................................................................................................... 3
2.3.
Minotaur IV Launch Vehicle .......................................................................................................... 4
2.3.1.
Stage 1, 2 and 3 Booster Assemblies....................................................................................... 4
2.3.2.
Stage 4 Booster/Avionics Assembly ......................................................................................... 5
2.3.2.1.
Avionics............................................................................................................................... 5
2.3.2.2.
Attitude Control Systems .................................................................................................... 5
2.3.2.3.
Telemetry Subsystem ......................................................................................................... 7
2.3.2.4.
Payload Fairing ................................................................................................................... 7
2.4.
Launch Support Equipment .......................................................................................................... 8
2.4.1. 3.
Transportable LSE Shelters ...................................................................................................... 8
GENERAL PERFORMANCE ............................................................................................................ 11 3.1.
Mission Profiles........................................................................................................................... 11
3.2.
Launch Sites ............................................................................................................................... 11
3.2.1.
Western Launch Sites ............................................................................................................. 11
3.2.2.
Eastern Launch Sites .............................................................................................................. 11
3.2.3.
Alternate Launch Sites ............................................................................................................ 11
3.3.
Performance Capability............................................................................................................... 11
3.4.
Injection Accuracy....................................................................................................................... 23
3.5.
Payload Deployment................................................................................................................... 23
3.6.
Payload Separation..................................................................................................................... 23
3.7.
Collision/Contamination Avoidance Maneuver ........................................................................... 23
4.
PAYLOAD ENVIRONMENT.............................................................................................................. 25 4.1.
Steady State and Transient Acceleration Loads......................................................................... 25
4.1.1.
Transient Loads....................................................................................................................... 26
4.1.2.
Steady-State Acceleration....................................................................................................... 26
4.2.
Payload Vibration Environment .................................................................................................. 26
4.2.1.
Random Vibration ................................................................................................................... 28
4.2.2.
Sine Vibration .......................................................................................................................... 28
4.3.
Payload Acoustic Environment ................................................................................................... 28
4.4.
Payload Shock Environment....................................................................................................... 28
4.5.
Payload Structural Integrity and Environments Verification........................................................ 32
4.5.1. 4.6.
Recommended Payload Testing and Analysis ....................................................................... 33 Thermal and Humidity Environments.......................................................................................... 33
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4.6.1.
Ground Operations.................................................................................................................. 33
4.6.2.
Powered Flight ........................................................................................................................ 34
4.6.3.
Nitrogen Purge (Non-Standard Service) ................................................................................. 35
4.7.
Payload Contamination Control .................................................................................................. 35
4.8.
Payload Electromagnetic Environment....................................................................................... 35
5.
PAYLOAD INTERFACES.................................................................................................................. 37 5.1.
Payload Fairing ........................................................................................................................... 37
5.1.1.
Payload Dynamic Design Envelope ........................................................................................ 37
5.1.2.
Payload Access Door.............................................................................................................. 37
5.2.
Payload Mechanical Interface and Separation System .............................................................. 37
5.2.1.
Standard Non-Separating Mechanical Interface ..................................................................... 38
5.2.2.
Orbital Supplied Mechanical Interface Control Drawing ......................................................... 38
5.3.
Payload Electrical Interfaces....................................................................................................... 38
5.3.1.
Payload Umbilical Interfaces................................................................................................... 38
5.3.2.
Payload Interface Circuitry ...................................................................................................... 40
5.3.3.
Payload Battery Charging ....................................................................................................... 40
5.3.4.
Payload Command and Control .............................................................................................. 40
5.3.5.
Pyrotechnic Initiation Signals .................................................................................................. 40
5.3.6.
Payload Telemetry .................................................................................................................. 40
5.3.7.
Payload Separation Monitor Loopbacks ................................................................................. 41
5.3.8.
Telemetry Interfaces ............................................................................................................... 41
5.3.9.
Non-Standard Electrical Interfaces ......................................................................................... 41
5.3.10.
Electrical Launch Support Equipment................................................................................... 41
5.4.
Payload Design Constraints........................................................................................................ 41
5.4.1.
Payload Center of Mass Constraints ...................................................................................... 41
5.4.2.
Final Mass Properties Accuracy.............................................................................................. 41
5.4.3.
Pre-Launch Electrical Constraints........................................................................................... 42
5.4.4.
Payload EMI/EMC Constraints................................................................................................ 42
5.4.5.
Payload Dynamic Frequencies ............................................................................................... 42
5.4.6.
Payload Propellant Slosh ........................................................................................................ 42
5.4.7.
Payload-Supplied Separation Systems................................................................................... 42
5.4.8.
System Safety Constraints...................................................................................................... 42
6.
MISSION INTEGRATION.................................................................................................................. 43 6.1.
Mission Management Approach ................................................................................................. 43
6.1.1.
RSLP Mission Responsibilities................................................................................................ 43
6.1.2.
Orbital Mission Responsibilities .............................................................................................. 43
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6.2.
Mission Planning and Development ........................................................................................... 44
6.3.
Mission Integration Process........................................................................................................ 44
6.3.1.
Integration Meetings................................................................................................................ 44
6.3.2.
Mission Design Reviews (MDR).............................................................................................. 45
6.3.3.
Readiness Reviews................................................................................................................. 45
6.4.
Documentation............................................................................................................................ 46
6.4.1.
Customer-Provided Documentation ........................................................................................ 46
6.4.1.1.
Payload Questionnaire...................................................................................................... 46
6.4.1.2.
Payload Mass Properties .................................................................................................. 46
6.4.1.3.
Payload Finite Element Model .......................................................................................... 46
6.4.1.4.
Payload Thermal Model for Integrated Thermal Analysis................................................. 46
6.4.1.5.
Payload Drawings ............................................................................................................. 46
6.4.1.6.
Program Requirements Document (PRD) Mission Specific Annex Inputs ....................... 46
6.4.1.6.1. 6.5.
Launch Operations Requirements (OR) Inputs .......................................................... 47
Safety .......................................................................................................................................... 47
6.5.1.
System Safety Requirements.................................................................................................. 47
6.5.2.
System Safety Documentation................................................................................................ 47
7.
GROUND AND LAUNCH OPERATIONS ......................................................................................... 49 7.1.
Minotaur IV/Payload Integration Overview ................................................................................. 49
7.2.
Ground And Launch Operations ................................................................................................. 49
7.2.1.
Launch Vehicle Integration...................................................................................................... 49
7.2.1.1.
Planning and Documentation............................................................................................ 49
7.2.1.2.
GCA/Orion 38 Integration and Test Activities ................................................................... 49
7.2.1.3.
PK Motor Integration and Test Activities........................................................................... 49
7.2.1.4.
Mission Simulation Tests .................................................................................................. 49
7.2.1.5.
Booster Assembly Stacking/Launch Pad Preparation ...................................................... 51
7.2.2.
Payload Processing/Integration .............................................................................................. 51
7.2.2.1.
Payload Propellant Loading .............................................................................................. 52
7.2.2.2.
Final Vehicle Integration and Test .................................................................................... 52
7.3.
Launch Operations...................................................................................................................... 53
7.3.1. 8.
Launch Control Organization .................................................................................................. 53
OPTIONAL ENHANCED CAPABILITIES.......................................................................................... 55 8.1.
Mechanical Interface and Separation System Enhancements ................................................... 55
8.1.1.
Separation Systems ................................................................................................................ 55
8.1.2.
Additional Fairing Access Doors ............................................................................................. 55
8.1.3.
Payload Isolation System........................................................................................................ 55
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Performance Enhancements ...................................................................................................... 56
8.2. 8.2.1.
Insertion Accuracy................................................................................................................... 56
8.2.2.
Star 48 Stage 4 ....................................................................................................................... 57
8.3.
Environmental Control Options ................................................................................................... 58
8.3.1.
Conditioned Air........................................................................................................................ 58
8.3.2.
Nitrogen Purge ........................................................................................................................ 58
8.3.3.
Enhanced Contamination Control ........................................................................................... 58
8.3.3.1.
High Cleanliness Integration Environment (Class 10K or 100K) ...................................... 59
8.3.3.2.
Fairing Surface Cleanliness Options ................................................................................ 59
8.3.3.3.
High Cleanliness Fairing Environment.............................................................................. 59
8.3.4.
Launch Pad Environmental Control ........................................................................................ 59
8.3.4.1. 8.4.
Booster Temperature Control ........................................................................................... 59
Enhanced Telemetry Options ..................................................................................................... 60
8.4.1.
Enhanced Telemetry Bandwidth ............................................................................................. 60
8.4.2.
Enhanced Telemetry Instrumentation ..................................................................................... 60
8.4.3.
Navigation Data....................................................................................................................... 60
8.5.
Shared Launch Accommodations ............................................................................................... 60 LIST OF FIGURES
Figure 2-1. OSP-2 Peacekeeper Space Lift Vehicle.................................................................................... 3 Figure 2-2. Minotaur IV Expanded View ...................................................................................................... 4 Figure 2-3. Orion 38 Stage 4 Motor ............................................................................................................. 6 Figure 2-4. Stage 4 Structures..................................................................................................................... 6 Figure 2-5. Existing 92 in. Taurus Fairing, Handling Fixtures, and Processes will be used for Minotaur IV....................................................................................................... 7 Figure 2-6. Functional Block Diagram of LSE.............................................................................................. 9 Figure 2-7. Portable Launch Support Structure Provide Optional Support From Austere Sites................ 10 Figure 3-1. Minotaur IV Mission Profile...................................................................................................... 12 Figure 3-2. Minotaur IV Launch Site Options............................................................................................. 13 Figure 3-3. Minotaur IV Performance Curves for VAFB Launches (Metric Units) ..................................... 14 Figure 3-4. Minotaur IV Performance Curves for VAFB Launches (English Units) ................................... 15 Figure 3-5. Minotaur IV Performance Curves for CCAFS Launches (Metric Units) .................................. 16 Figure 3-6. Minotaur IV Performance Curves for CCAFS Launches (English Units) ................................ 17 Figure 3-7. Minotaur IV Performance Curves for Kodiak, Alaska Launches (Metric Units)....................... 18 Figure 3-8. Minotaur IV Performance Curves for Kodiak, Alaska Launches (English Units)..................... 19 Figure 3-9. Minotaur IV Performance Curves for Wallops, Virginia Launches (Metric Units).................... 20 Figure 3-10. Minotaur IV Performance Curves for Wallops, Virginia Launches (English Units)................ 21 Release 1.1
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Figure 3-11. Stage Impact Points for VAFB and CCAFS Launches.......................................................... 22 Figure 3-12. Minotaur IV Injection Accuracy .............................................................................................. 24 Figure 3-13. Typical Pre-Separation Payload Pointing and Spin Rate Accuracies .................................. 24 Figure 4-1. Phasing of Dynamic Loading Events....................................................................................... 25 Figure 4-2. Payload CG Net Transient Lateral Acceleration at Stage 2 Ignition with a Typical Separation System................................................................................................................... 27 Figure 4-3. Minotaur IV Nominal Maximum Axial Acceleration as a Function of Payload Mass ............... 27 Figure 4-4. Minotaur IV Payload Random Vibration Environment ............................................................. 29 Figure 4-5. Minotaur IV Payload Sine Vibration Environment ................................................................... 29 Figure 4-6. Minotaur IV Payload Acoustic Maximum Predicted Environment (MPE) ................................ 30 Figure 4-7. Minotaur IV Payload Shock Maximum Predicted Environment (MPE) – Launch Vehicle to Payload .................................................................................................................................... 31 Figure 4-8. Payload Shock Environment – Payload to Launch Vehicle .................................................... 32 Figure 4-9. Factors of Safety Payload Design and Test ........................................................................... 33 Figure 4-10. Recommended Payload Testing Requirements.................................................................... 33 Figure 4-11. Payload Thermal and Humidity Environment ........................................................................ 34 Figure 4-12. Minotaur IV Launch Vehicle RF Emitters and Receivers ...................................................... 36 Figure 5-1. Standard 92 in. Fairing Envelope ............................................................................................ 37 Figure 5-2. Standard, Non-separating Payload Mechanical Interface....................................................... 39 Figure 5-3. Payload Electrical Interface Block Diagram, With No Orbital Supplied Separation System .. 39 Figure 5-4. Payload 1:1 Umbilical Pin Outs ............................................................................................... 39 Figure 5-5. Minotaur IV Payload Electrical Interface Block Diagram ......................................................... 40 Figure 5-6. Payload Mass Properties Measurement Tolerance ................................................................ 42 Figure 6-1. Typical Integrated OSP Organizational Structure ................................................................... 43 Figure 6-2. Typical Mission Integration Schedule...................................................................................... 45 Figure 7-1. Hardware Flow – Factory to Launch Site ................................................................................ 50 Figure 7-2. SLV Processing Flow .............................................................................................................. 51 Figure 7-3. Minotaur IV Processing Flow................................................................................................... 52 Figure 7-4. Minotaur IV Upper Stack Assembly will be Vertically Integrated to Minotaur IV Booster Assembly in a Similar Manner to Taurus Upper Stack ............................................................ 53 Figure 8-1. 38-in. Separation System Option............................................................................................. 56 Figure 8-2. Soft Ride Payload Isolation System as Integrated on Minotaur LV......................................... 57 Figure 8-3. Hydrazine Auxiliary Propulsion System (HAPS) Used to Provide Insertion Accuracy............ 57 Figure 8-4. Orion 38 Stage 4 Motor can be Replaced with a Star-48 to Provide Increased Performance 58 Figure 8-5. Mobile Scaffolding for Environmental Control Demonstrated on Minotaur Missions .............. 59 Figure 8-6. Modular Minotaur IV Structural Design Easily Accommodates Multiple Payloads ................. 60
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Table of Contents LIST OF TABLES PAGE
TABLE 4-1. PAYLOAD CG PARAMETRIC DESIGN LIMIT LOADS ......................................................... 26
LIST OF APPENDICES A. PAYLOAD QUESTIONNAIRE..............................................................................................................A-1
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Minotaur IV User’s Guide ACS A/D AADC AC AFRL BER C/CAM CCAFS CDR CG CLA CLF CMP DPAF ECS EELV EGSE EMC EME EMI FLSA FTS GCA GFE GFP GPB GPS GTO HAPS HEPA HVAC ICD LCR LER LEV LOCC LRR LSE LSV LV MAC MAC
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Attitude Control System Arm/Disarm Alaska Aerospace Development Corporation Air Conditioning Air Force Research Laboratory Bit Error Rate Collision/Contamination Avoidance Maneuver Cape Canaveral Air Force Station Critical Design Review center-of-gravity Coupled Loads Analysis Commercial Launch Facility Critical Measurements Program Dual Payload Attach Fitting Environmental Control System Evolved Expendable Launch Vehicle Electrical Ground Support Equipment Electromagnetic Compatibility Electromagnetic Environment Electromagnetic Interference Florida Spaceport Authority Flight Termination System Guidance and Control Assembly Government Furnished Equipment Government Furnished Property GPS Position Beacon Global Positioning System Geosynchronous Transfer Orbit Hydrazine Auxiliary Propulsion System High Efficiency Particulate Air Heating, Ventilation and Air Conditioning Interface Control Drawing Launch Control Room Launch Equipment Room Launch Equipment Van Launch Operations Control Center Launch Readiness Review Launch Support Equipment Launch Support Van Launch Vehicle Motor Adapter Cone Modified Adapter Cone
MACH MARS MDR MIWG MRD MRR NTW ODM OR OSP-2 PAF PAM PBCM PCF PCM PDR PID PK POC PPF PRD PWB QES RF RSLP RTS RWG SEA SEB SES SLV SRSS SSI START TPS TVAs TVC USAF VAFB VPF WFF WP
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Modular Avionics Control Hardware Mid-Atlantic Regional Spaceport Mission Design Review Mission Integration Working Groups Mission Requirements Document Mission Readiness Review Navy Theater Wide Ordnance Driver Modules Operations Requirements Orbital Suborbital Program ll Payload Attach Fitting Payload Adapter Module PK Booster Control Module Programmable Clock Filter Pulse Code Modulation Preliminary Design Review Proportional-Integral-Derivative Peacekeeper Point of Contact Payload Processing Facility Program Requirements Document Printed Wiring Board Quick Erect Scaffold Radio Frequency Rocket Systems Launch Program Range Tracking System Range Working Groups Statistical Energy Analysis Support Equipment Building Saab Ericson Space Space Launch Vehicle Softride for Small Satellites Spaceport Systems International Strategic Arms Reduction Treaty Thermal Protection System Thrust Vector Actuators Thrust Vector Control United States Air Force Vandenberg Air Force Base Vehicle Processing Facility Wallops Flight Facility Work Package
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Section 1.0 - Introduction
1. INTRODUCTION This User’s Guide is intended to familiarize payload mission planners with the capabilities of the Orbital Suborbital Program ll (OSP-2) Minotaur IV Space Launch Vehicle (SLV) launch service. The information provided in this user’s guide is for initial planning purposes only. Information for development/design is determined through mission specific engineering analyses. The results of these analyses are documented in a mission-specific Interface Control Document (ICD) for the payloader organization to use in their development/design process. This document provides an overview of the Minotaur IV system design and a description of the services provided to our customers. Minotaur IV offers a variety of enhanced options to allow the maximum flexibility in satisfying the objectives of single or multiple payloads.
By adopting the austere launch site concepts developed for Taurus, the Minotaur IV 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 Force Station (CCAFS), Wallops Flight Facility (WFF), and Kodiak Island, Alaska. This User’s Guide describes Minotaur IVunique integration and test approaches (including the typical operational timeline for payload integration with the Minotaur IV vehicle) and the ground support equipment that will be used to conduct Minotaur IV operations.
The primary mission of the Minotaur IV is to provide low cost, high reliability launch services to government-sponsored payloads. Minotaur IV accomplishes this using flight-proven components with a significant flight heritage such as surplus Peacekeeper boosters, the Taurus Fairing and Attitude Control System, and a mix of Minotaur I, Pegasus, Taurus, and other orbital standard avionics, all with a proven, successful track record. The philosophy of placing mission success as the highest priority is reflected in the success and accuracy of all OSP missions to date. The Minotaur IV launch vehicle system is composed of a flight vehicle and ground support equipment. Each element of the Minotaur IV system has been developed to simplify the mission design and payload integration process and to provide safe, reliable space launch services. This User’s Guide describes the basic elements of the Minotaur IV system as well as optional services that are available. In addition, this document provides general vehicle performance, defines payload accommodations and environments, and outlines the Minotaur IV mission integration process.
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Section 2.0 – Minotaur IV Launch Service
2. MINOTAUR IV LAUNCH SERVICE 2.1. Minotaur IV Launch System Overview The Minotaur IV (Figure 2-1) mission is to provide a cost effective, reliable and flexible means of placing satellites into orbit. The launch vehicle developer and manufacturer is Orbital, under the Orbital Suborbital Program 2 (OSP-2) contract for the U.S. Air Force. An overview of the system and available launch services is provided within this section, with specific elements covered in greater detail in the subsequent sections of this User’s Guide.
IV draws on the successful heritage of four launch vehicles: Orbital’s Pegasus, Taurus and Minotaur I space launch vehicles and the Peacekeeper system of the USAF. Minotaur IV’s avionics are derived from the Pegasus and Taurus systems, providing a combined total of more than 35 successful space launch missions. Orbital’s efforts have enhanced or updated Pegasus, Taurus and Minotaur avionics components to meet the payload-support requirements of the OSP-2 program. Combining these improved subsystems with the long successful history of the Peacekeeper boosters has resulted in a simple, robust, self-contained launch system to support government-sponsored small satellite launches. The Minotaur IV system also includes a complete set of transportable Launch Support Equipment (LSE) designed to allow Minotaur IV to be operated as a self-contained satellite delivery system. To accomplish this goal, the Electrical Ground Support Equipment (EGSE) has been developed to be portable and adaptable to varying levels of infrastructure. While the Minotaur IV system is capable of self-contained operation using portable vans to house the EGSE, it is typically launched from an established range where the EGSE can be housed in available, permanent structures or facilities.
Figure 2-1. OSP-2 Peacekeeper Space Lift Vehicle Minotaur IV has been designed to meet the needs of U.S. Government-sponsored customers at a lower cost than commercially available alternatives by the use of surplus Peacekeeper boosters. The requirements of the OSP-2 program stress system reliability, transportability, and operation from multiple launch sites. Minotaur
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The vehicle and LSE are designed to be capable of launch from any of the four commercial Spaceports (Alaska, California, Florida, and Mid-Atlantic), as well as from existing U.S. Government facilities at VAFB and CCAFS. The Launch Control Room (LCR) serves as the control center for conducting a Minotaur IV launch and includes consoles for Orbital, range safety, and limited customer personnel. Further description of the Launch Support Equipment is provided in Section 2.4. 2.2. Minotaur IV Launch Service The Minotaur IV Launch Service is provided through the combined efforts of the USAF and Orbital, along with associate contractors including
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Northrop Grumman and Commercial Spaceports. The primary customer interface will be with the USAF Space and Missile Systems Center, Detachment 12, Rocket Systems Launch Program, designated RSLP. Orbital is the launch vehicle provider. For brevity, this integrated team effort will be referred to as “OSP”. Where interfaces are directed toward a particular member of the team, they will be referred to directly (i.e., Orbital or RSLP). OSP provides all of the necessary hardware, software and services to integrate, test and launch a satellite into its prescribed orbit. In addition, OSP will complete all the required agreements, licenses and documentation to successfully conduct Minotaur IV operations. The Minotaur IV mission integration process completely identifies, documents, and verifies all spacecraft and mission requirements. This provides a solid basis for initiating and streamlining the integration process for future Minotaur IV customers. 2.3. Minotaur IV Launch Vehicle The Minotaur IV vehicle, shown in expanded view in Figure 2-2, is a four-stage, inertially guided, all solid propellant ground launched vehicle. Conservative design margins, state-ofthe-art structural systems, a modular avionics architecture, and simplified integration and test capability, yield a robust, highly reliable launch vehicle design. In addition, Minotaur IV payload accommodations and interfaces have been designed to satisfy a wide range of potential payload requirements. 2.3.1. Stage 1, 2 and 3 Booster Assemblies The first three stages of the Minotaur IV consists of the refurbished Government Furnished Equipment (GFE) Peacekeeper Stages 1, 2 and 3. These booster assemblies are used as provided by the Government, requiring no modification or additional components. They have extensive flight history, including 51 Peacekeeper launches and three Stage 1 launches on Taurus with no motor related failures. All three stages are solid-
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Figure 2-2. Minotaur IV Expanded View propellant motors and utilize a movable nozzle controlled by a Thrust Vector Actuator (TVA) system for three-axis attitude control. The first
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stage consists of a Thiokol motor that provides 500,000 lbf (2224 kN) of thrust. The second stage motor is an Aerojet motor with a moveable nozzle contoured with an extendable exit cone. It provides an average thrust of 275,000 lbf (1223 kN). The third stage is a Hercules motor that provides 65,000 lbf (289 kN) of thrust and also features an extendable exit cone similar to Stage 2. 2.3.2. Stage 4 Booster/Avionics Assembly The Minotaur IV Stage 4 motor is the Orion 38 design (Figure 2-3). This motor was originally developed for Orbital’s Pegasus program and is used on the Minotaur I launch vehicle, as well as other Orbital launch vehicles. Common design features, materials and production techniques are applied to the motors to maximize reliability and production efficiency. The Orion 38 motor provides the velocity needed for orbit insertion for the SLV, in the same manner as it is used on the Minotaur. This motor features state-of-the-art design and materials with a successful flight heritage. It is currently in production and is actively flying payloads into space, with 37 fully successful flights to date and one static test. Three substructures are utilized to accommodate the Orion 38 Stage 4 motor and attach it to the Stage 1-3 PK booster assembly. These are a Payload Adapter Module (PAM) with 62.01-inch diameter payload interface, a 38-inch diameter motor adapter cone and a GCA +3/4 interstage. These structures were adapted from similar Taurus hardware designs and are shown in Figure 2-4. 2.3.2.1. Avionics The Minotaur IV avionics system has heritage to the Minotaur I, OSP TLV, as well as Pegasus and Taurus designs. The flight computer is a 32bit multiprocessor architecture. It provides communication with vehicle subsystems, the LSE, and the payload, if required, utilizing standard RS422 serial links and discrete I/O. The avionics system design incorporates Orbital’s innovative,
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flight-proven Modular Avionics Control Hardware (MACH). Standardized, function-specific modules are combined in stacks to meet vehicle-specific requirements. The functional modules from which the MACH stacks are created include power transfer, ordnance initiation, booster interface, communication, and telemetry processing. Orbital has designed, tested, and flown a variety of MACH modules, which provide an array of functional capability and flexibility. MACH has exhibited 100% reliability on all flights including Minotaur and TLV flights and several of Orbital’s suborbital launch vehicles including Navy Theater Wide (NTW), Critical Measurements Program (CMP), and STORM. 2.3.2.2. Attitude Control Systems The Minotaur IV Attitude Control System (ACS) provides attitude control throughout boosted flight and coast phases. Stages 1, 2 and 3 utilize the PK Thrust Vector Control (TVC) systems. The PK Booster Control Module (PBCM) links the flight computer actuator commands to the individual Stage 1, 2, and 3 Thrust Vector Actuators (TVAs). Stage 4 utilizes the same TVC system used by the Pegasus, Taurus and Minotaur vehicles which combines single-nozzle electromechanical TVC for pitch and yaw control with a three-axis, cold-gas attitude control system resident in the avionics section providing roll control. Attitude control is achieved using a three-axis autopilot that employs Proportional-IntegralDerivative (PID) control. Stages 1, 2 and 3 fly a pre-programmed attitude profile based on trajectory design and optimization. Stage 4 uses a set of pre-programmed orbital parameters to place the vehicle on a trajectory toward the intended insertion apse. The extended coast between Stages 3 and 4 is used to orient the vehicle to the appropriate attitude for Stage 4 ignition based upon a set of pre-programmed orbital parameters and the measured performance of the first three stages. Stage 4 utilizes energy management to place the vehicle into the proper orbit. After the
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Figure 2-3. Orion 38 Stage 4 Motor
Figure 2-4. Stage 4 Structures
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final boost phase, the three-axis cold-gas attitude control system is used to orient the vehicle for spacecraft separation, contamination and collision avoidance and downrange downlink maneuvers. The autopilot design is modular, so additional payload requirements such as rate control or celestial pointing can be accommodated with minimal additional development. 2.3.2.3. Telemetry Subsystem The Minotaur IV telemetry subsystem provides real-time health and status data of the vehicle avionics system, as well as key information regarding the position, performance and environment of the Minotaur IV vehicle. This data may be used by Orbital and the range safety personnel to evaluate system performance. The Minotaur IV baseline telemetry subsystem provides a number of dedicated payload discrete (bi-level) and analog telemetry monitors through dedicated channels in the SLV encoder. The Minotaur IV telemetry system provides a baseline 1 Mbps data rate (both payload and Minotaur IV telemetry). However, the output data rate is selectable from 2.441 kbps to 10 Mbps to allow flexibility in supporting evolving mission requirements, as limited by link margin and Bit Error Rate (BER) requirements. The telemetry subsystem nominally utilizes Pulse Code Modulation (PCM) with a RNRZ-L format. However other types of data formats, including NRZ-L, S, M, and Bi-phase may be implemented if required, in order to accommodate launch range limitations.
control during ground handling, integration operations and flight. The fairing is a bi-conic design made of graphite/epoxy face sheets with an aluminum honeycomb core. The fairing provides for low payload contamination through prudent design and selection of low contamination materials and processes. Acoustic blankets and internal air conditioning ducts are available to provide more benign payload environments. Air conditioning will keep the payload environment to a specified temperature between 60 to 120 °F dependent upon requirements. The two halves of the fairing are structurally joined along their longitudinal interface using Orbital’s low contamination frangible joint system. An additional circumferential frangible joint at the base of the fairing supports the fairing loads. At separation, a gas pressurization system is activated to pressurize the fairing deployment thrusters. The fairing halves then rotate about external hinges that control the fairing deployment to ensure that payload and launch vehicle clearances are maintained. All elements of the deployment system have been demonstrated through test to comply with stringent contamination requirements. Options for payload access doors and enhanced cleanliness are available. Further
Minotaur IV telemetry is subject to the provisions of the Strategic Arms Reduction Treaty (START). START treaty provisions require that certain Minotaur IV telemetry be unencrypted and provided to the START treaty office for dissemination to the signatories of the treaty. 2.3.2.4. Payload Fairing Orbital’s flight-proven Taurus 92-inch diameter payload fairing (Figure 2-5) is used to encapsulate the payload, provide protection and contamination
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Figure 2-5. Existing 92 in. Taurus Fairing, Handling Fixtures, and Processes will be used for Minotaur IV
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details on the baseline fairing are included in Section 5.1. With the addition of a structural adapter, the fairing can accommodate multiple payloads. This feature, described in more detail in Section 8.5 of this User’s Guide, permits two or more smaller payloads to share the cost of a Minotaur launch, resulting in a lower launch cost for each as compared to other launch options. OSP has access to several Multiple Payload Adaptor (MPA) designs that allow for a cost sharing benefit to programs with excess payload and/or mass capability. 2.4. Launch Support Equipment The Minotaur IV LSE is designed to be readily adaptable to varying launch site configurations with minimal unique infrastructure required. The EGSE consists of readily transportable consoles that can be housed in various facility configurations depending on the launch site infrastructure. The EGSE is composed of three primary functional elements: Launch Control, Vehicle Interface, and Telemetry Data Reduction. The Launch Control consoles are located in a Launch Control Room (LCR), or mobile launch equipment van depending on available launch site accommodations. The Vehicle Interface EGSE is located in a permanent structure, typically called a Support Equipment Building (SEB) or Launch Equipment Room (LER). Fiber optic connections from the Launch Control to the Vehicle Interface consoles are used for efficient, high bandwidth communications and eliminates the need for copper wire. The Vehicle Interface racks provide the junction from fiber optic cables to the copper cabling interfacing with the vehicle. Figure 2-6 depicts the functional block diagram of the LSE. The LCR serves as the control center during the launch countdown. The number of personnel that can be accommodated are dependent on the launch site facilities. At a minimum, the LCR will accommodate Orbital personnel controlling the vehicle, two Range Safety representatives (ground and flight safety), and the Air Force Mission Manager. Mission-unique, customer-supplied
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payload consoles and equipment can be supported in the LCR and SEB, within the constraints of the launch site facilities or temporary structure facilities. Interface to the payload through the Minotaur IV payload umbilicals and landlines provides the capability for direct monitoring of payload functions. Payload personnel accommodations will be handled on a mission-specific basis. 2.4.1. Transportable LSE Shelters In order to perform mission operations from alternative, austere launch sites, Orbital can provide self contained, transportable shelters for the Launch Support Van (LSV) and Launch Equipment Van (LEV) as an unpriced option. These shelters are the same approach and design used on all six of Orbital’s Taurus launches (Figure 2-7). The OSP-2 Ground Support Consoles have been intentionally made modular and portable to allow their use in these accommodations. The LSV consists of a shelter which is located at a Range Safety-approved man-safe distance from the launch site. The LSV contains the vehicle control and telemetry monitor consoles described in Section 2.4. Sufficient space is available for additional equipment racks depending on Government and/or Payload requirements. The LEV consists of an 8 foot x 20 foot shelter which is located near the launch stool and is unmanned during launch. This shelter contains the vehicle interface racks described in Section 2.4. The LEV has sufficient room available for payload power supplies and interface electronics. Both shelters are designed for shipping and transportation with exterior tiedown and anchor locations used to facilitate the loading and unloading operations. The shelters can be delivered to any level location and be set up within hours. The transportable support console design also allows for the LSE to be moved into a fixed blockhouse and LER if required.
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Figure 2-6. Functional Block Diagram of LSE
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Figure 2-7. Portable Launch Support Structure Provide Optional Support From Austere Sites
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3. GENERAL PERFORMANCE
3.2.2. Eastern Launch Sites For Easterly launch azimuths to achieve orbital inclinations between 28.5° and 60°, Minotaur IV can be launched from facilities at Cape Canaveral, FL or Wallops Island, VA. Launches from Florida will nominally use the Florida Spaceport Authority (FLSA) launch facilities at LC-46 on CCAFS. These will be typically for inclinations from 28.5° to 40°, although inclinations above 35° may have reduced performance due to the need for a trajectory dogleg. Inclinations below 28.5° are feasible, albeit with doglegs and altitude constraints due to range safety considerations. The Mid-Atlantic Regional Spaceport (MARS) facilities at the WFF may be used for inclinations from 30° to 60°. Some inclinations may have reduced performance due to range safety considerations and will need to be evaluated on a case-by-case missionspecific basis.
3.1. Mission Profiles Minotaur IV can attain a range of posigrade and retrograde inclinations through the choice of launch sites made available by the readily adaptable nature of the Minotaur IV launch system. A typical mission profile to a sunsynchronous orbit is shown in Figure 3-1. All performance parameters presented herein are typical for most expected payloads. However, performance may vary depending on unique payload or mission characteristics. Specific requirements for a particular mission should be coordinated with OSP. Once a mission is formally initiated, the requirements will be documented in the Mission Requirements Document (MRD). Further detail will be captured in the Payload-toLaunch Vehicle Interface Control Drawing (ICD). 3.2. Launch Sites Depending on the specific mission and range safety requirements, Minotaur IV can operate from several East and West Coast launch sites, illustrated in Figure 3-2. Specific performance parameters are presented in Section 3.3. Facilities used for OSP Minotaur and Taurus launches are generally compatible with Minotaur IV operations. 3.2.1. Western Launch Sites For missions requiring high inclination orbits (greater than 60°), launches can be conducted from facilities at VAFB or Kodiak Island, AK. Both facilities can accommodate inclinations from 60° to 120°, although inclinations below 72° from VAFB require an out-of-plane dogleg, thereby reducing payload capability. As with the initial Minotaur missions, Minotaur IV can be launched from the California Spaceport facility, Space Launch Complex 8 (SLC-8) operated by Spaceport Systems International (SSI), on South VAFB. The launch facility at Kodiak Island, operated by the Alaska Aerospace Development Corporation (AADC) has been used for both orbital and suborbital launches.
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3.2.3. Alternate Launch Sites Other launch facilities can be readily used given the flexibility designed into the Minotaur IV vehicle, ground support equipment, and the various interfaces. Orbital has experience launching vehicles from a variety of sites around the world. To meet the requirements of performing mission operations from alternative, austere launch sites, Orbital can provide self contained, transportable shelters as described in section 2.4.1. 3.3. Performance Capability Minotaur IV performance curves for circular and elliptical orbits of various altitudes and inclinations are detailed in Figure 3-3 through Figure 3-10 for launches from all four Spaceports in both metric and English units. These performance curves provide the total mass above the standard, non-separating interface. The mass of any Payload Attach Fitting (PAF) or separation system is to be accounted for in the payload mass allocation. Figure 3-11 illustrates the stage vacuum impact points for launch trajectories from VAFB and CCAFS.
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Figure 3-1. Minotaur IV Mission Profile
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Figure 3-2. Minotaur IV Launch Site Options
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Figure 3-3. Minotaur IV Performance Curves for VAFB Launches (Metric Units)
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Figure 3-4. Minotaur IV Performance Curves for VAFB Launches (English Units)
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Figure 3-5. Minotaur IV Performance Curves for CCAFS Launches (Metric Units)
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Figure 3-6. Minotaur IV Performance Curves for CCAFS Launches (English Units)
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Figure 3-7. Minotaur IV Performance Curves for Kodiak, Alaska Launches (Metric Units)
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Figure 3-8. Minotaur IV Performance Curves for Kodiak, Alaska Launches (English Units)
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Figure 3-9. Minotaur IV Performance Curves for Wallops, Virginia Launches (Metric Units)
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Figure 3-10. Minotaur IV Performance Curves for Wallops, Virginia Launches (English Units)
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Figure 3-11. Stage Impact Points for VAFB and CCAFS Launches
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3.4. Injection Accuracy Minotaur IV injection accuracy is summarized in Figure 3-12. Better accuracy can be provided dependent on specific mission characteristics. For example, heavier payloads will typically have better insertion accuracy, as will higher orbits. An enhanced option for increased insertion accuracy is also available (Section 8.2.1). It utilizes the flight-proven Hydrazine Auxiliary Propulsion System (HAPS) developed on the Pegasus program. 3.5. Payload Deployment Following orbit insertion, the Minotaur IV Stage 4 avionics subsystem can execute a series of ACS maneuvers to provide the desired initial payload attitude prior to separation. This capability may also be used to incrementally reorient Stage 4 for the deployment of multiple spacecraft with independent attitude requirements. Either an inertially-fixed or spin-stabilized attitude may be specified by the customer. The maximum spin rate for a specific mission depends upon the spin axis moment of inertia of the payload and the amount of ACS propellant needed for other attitude maneuvers. Figure 3-13 provides the typical payload pointing and spin rate accuracies. 3.6. Payload Separation Payload separation dynamics are highly dependent on the mass properties of the payload and the particular separation system utilized. The primary parameters to be considered are payload tip-off and the overall separation velocity.
Separation velocities are driven by the need to prevent recontact between the payload and the Minotaur IV 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 IV 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 IV 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 “crabwalk” maneuver to impart a small amount of delta velocity, increasing the separation between the payload and the fourth stage of the Minotaur IV. Following the completion of the C/CAM maneuver as described above and any remaining maneuvers, such as downlinking of delayed telemetry data (per START treaty provisions), the ACS valves are opened and the remaining ACS nitrogen propellant is expelled.
Payload tip-off refers to the angular velocity imparted to the payload upon separation due to payload center-of-gravity (CG) offsets and an uneven distribution of torques and forces. If an optional Orbital-supplied Marmon Clamp-band separation system is used, payload tip-off rates are generally under 5°/sec per axis. Separation system options are discussed further in Section 8.1.1. Orbital performs a mission-specific tip-off analysis for each payload.
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Error Type
Section 3.0 – General Performance
Tolerance (Worst Case)
Error Source
Altitude (Insertion Apse)
±10 nmi (18.5 km)
Stage 4 motor performance uncertainty and guidance algorithm uncertainty
Altitude (Non-Insertion Apse)
±50 nmi (92.6 km)
Stage 4 motor performance and guidance algorithm uncertainty and navigation (INS) error
Altitude (Mean)
±30 nmi (55.6 km)
Stage 4 motor performance and guidance algorithm uncertainty and navigation (INS) error
Inclination
±0.2°
Guidance algorithm uncertainty and navigation (INS) error
Figure 3-12. Minotaur IV Injection Accuracy
Figure 3-13. Typical Pre-Separation Payload Pointing and Spin Rate Accuracies
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Section 4.0 – Payload Environment
4. PAYLOAD ENVIRONMENT
events such as motor ignition, stage separation, and transonic crossover.
CAUTION The environmental design and test criteria presented have been derived using measured data obtained from previous Taurus, Minotaur I, and Peacekeeper missions, motor static fire tests, other system development tests and analyses. The predicted levels presented are intended to be representative of a standard mission. Satellite mass, geometry and structural components vary greatly and will result in significant differences from mission to mission.
The predicted environments provided in this user's guide are for initial planning purposes only. Environments presented here bound typical mission parameters, but should not be used in lieu of mission-specific analyses. Mission-specific levels are provided as a standard service and documented or referenced in the mission ICD. This section provides details of the predicted environmental conditions that the payload will experience during Minotaur IV ground operations, powered flight, and launch system on-orbit operations. The predicted environments provided in this user’s guide are for initial planning purposes only. Minotaur IV ground operations include payload integration and encapsulation within the fairing, subsequent transportation to the launch site and final vehicle integration activities. Powered flight begins at Stage 1 ignition and ends at Stage 4 burnout. Minotaur IV on-orbit operations begin after Stage 4 burnout and end following payload separation. To more accurately define simultaneous loading and environmental conditions, the powered flight portion of the mission is further subdivided into smaller time segments bounded by critical, transient flight
Dynamic loading events that occur throughout various portions of the flight include steady-state acceleration, transient low frequency acceleration, acoustic impingement, random vibration, and pyrotechnic shock events. Figure 4-1 identifies the time phasing of these dynamic loading events and environments and their significance. Pyroshock events are not indicated in this figure, as they do not occur simultaneous with any other significant dynamic loading events. In addition, dynamic loading associated with S4 ignition is insignificant. 4.1. Steady State and Transient Acceleration Loads Design limit load factors due to the combined effects of steady state and low frequency transient accelerations are largely governed by payload characteristics. A mission-specific Coupled Loads Analysis (CLA) will be performed, with customer
Figure 4-1. Phasing of Dynamic Loading Events Release 1.1
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provided finite element models of the payload, in order to provide precise load predictions. Results will be referenced in the mission specific ICD. For preliminary design purposes, Orbital can provide initial Center-of-Gravity (CG) netloads given a payload’s mass properties, CG location and bending frequencies. Design limit loads due to both transient and steady-state accelerations are presented in Table 4-1 for select payload masses.
vehicle model. A Monte-Carlo analysis is performed to determine variations in vehicle acceleration due to changes in winds, motor performance and aerodynamics. The steady-state accelerations must be added to transient accelerations from the CLA as indicated in Figure 4-1 to determine the total payload acceleration. Steady-state accelerations are typically 8-Gs axial and 0.5-Gs lateral.
4.1.1. Transient Loads Transient loads account for approximately 30% of the total vehicle load with the remainder due to steady wind loads. Typical acceleration values at the payload interface are 2-Gs lateral and 4-Gs axial depending on the load case. Transient lateral accelerations at Stage 2 ignition are defined as a function of payload mass in Figure 4-2. Preliminary and final CLAs will be performed for each Minotaur IV payload. The payload finite element model is coupled to the vehicle model. Forcing functions have been developed for different flight events or load cases. Load cases include liftoff, transonic, max q and stage and shroud separation.
During powered flight, the maximum steady state accelerations are dependent on the payload mass. The maximum level occurs during Stage 3 burn. Figure 4-3 depicts the axial acceleration at burnout for Stages 3 and 4 as a function of payload mass.
4.1.2. Steady-State Acceleration The steady-state vehicle accelerations are determined from the vehicle rigid body analysis. Drag, wind and motor thrust are applied to a
4.2.
Payload Vibration Environment The Minotaur IV payload vibration environments are low frequency random and sinusoidal vibrations created by the solid rocket motor combustion processes and transmitted through the launch vehicle structure. Additionally, higher frequency aeroacoustics energy is created by air flow over the surface of the vehicle. Some of this aeroacoustic energy is transmitted via the launch vehicle structure to the payload. However, a majority of the aeroacoustic energy is transmitted to the payload directly as acoustic energy through the fairing.
TABLE 4-1. PAYLOAD CG PARAMETRIC DESIGN LIMIT LOADS Payload Mass 1600 lbm (725.7 kg) Axial (G) max/min Liftoff
3.83/0.27
2400 lb (1089 kg)
Lateral (G)
Axial (G)
0.62
3.93/0.15
max/min
3200 lb (1452 kg)
Lateral (G)
Axial (G)
0.46
3.90/0.16
max/min
4000lb (8141 kg)
Lateral (G)
Axial (G)
0.41
4.01/0.12
max/min
Lateral (G) 0.37
Pre-Transonic Resonant Burn
5.05/0.83
0.02
3.46/2.29
0.00
3.44/2.31
0.00
3.73/2.08
0.00
Transonic
5.13/1.52
1.23
3.95/2.71
0.97
3.89/2.75
0.90
4.07/2.49
0.89
Supersonic
3.41/3.40
1.96
3.38/3.38
1.61
3.36/3.36
1.40
3.34/3.34
1.26
Stage 2 Ignition
3.93/-0.35
4.05
3.95/-0.03
2.89
3.83/-0.02
2.74
3.66/0.01
2.13
Stage 3 Ignition
6.79/0.00
0.78
6.45/0.00
0.59
6.22/0.00
0.49
5.90/0.00
0.40
Stage 3 Burnout
See Figure 4-3
TBS
See Figure 4-3
TBS
See Figure 4-3
TBS
See Figure 4-3
TBS
Stage 4 Burnout
See Figure 4-3
TBS
See Figure 4-3
TBS
See Figure 4-3
TBS
See Figure 4-3
TBS
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Figure 4-2. Payload CG Net Transient Lateral Acceleration at Stage 2 Ignition with a Typical Separation System
Figure 4-3. Minotaur IV Nominal Maximum Axial Acceleration as a Function of Payload Mass
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4.2.1. Random Vibration Random vibration environments are produced both by the structurally transmitted vibrations through mechanical payload interface as well as from the acoustic energy directly through the fairing (Figure 4-4). However, the higher frequency aeroacoustic-induced random vibration levels are hard to accurately replicate for spacecraft of the Minotaur IV class. Testing input levels are typically hard to accurately specify because the response at the LV-to-spacecraft interface is strongly dependent on the unique spacecraft dynamics, including its response to the acoustic field. Therefore, structure-born random vibration environments are only defined herein for frequencies up to 250 Hz with the recommendation that high frequency environments be based on the acoustic levels defined in Section 4.3. The acoustic environment is defined starting at 20 Hz, allowing both environments to be evaluated in the overlapping region. The structure-borne low frequency random vibration environment for Minotaur IV is a flat 0.002 g2/Hz from 20 to 250 Hz, producing an overall level of 0.68 gRMS (Figure 4-4). This level is based on measured data from multiple applicable Orbital launch vehicles. Vibration produced by the PK Stage 1 motor was analyzed based on data from three Taurus flights, which utilized the PK Stage 1 motor. The random vibration levels also cover the Orion 38 motor burn observed on past Minotaur, Pegasus, and Taurus flights. Detailed evaluation of the Stage 2 and 3 motor burn is still pending, but prior experience indicates that their levels will be relatively benign for the payload. The Minotaur IV structure-borne vibration environment is enveloped by the MILSTD-1540E levels. OSP recommends that the 1540E level be used as guidance for minimum levels for payload testing. 4.2.2. Sine Vibration A known resonant burn characteristic of the PK Stage 1 motor creates a sine vibration requirement of Minotaur IV payloads. Orbital has Release 1.1
reviewed test data from more than 30 static fire tests and three Taurus flights. The sine vibration level depends on payload weight and stiffness. Orbital has developed forcing function to simulate the Stage 1 motor resonance. The forcing functions are used in a NASTRAN simulation to predict peak sine vibration levels at different vehicle locations. The simulation has been validated using flight data. The sine vibration varies between 45 and 75 Hz. The NASTRAN simulation of the PK Stage 1 resonant burn is analyzed for each payload as part of the CLA discussed in Section 4.1. The results of this mission-specific CLA are then used to refine the SLV vehicle and payload sine vibration levels. A parametric analysis of hypothetical payloads has been conducted to provide guidance on expected acceleration levels. The resulting levels are shown in Figure 4-5. 4.3.
Payload Acoustic Environment The acoustic environments to which the spacecraft will be exposed have been defined based on measured acoustic data from previous Taurus flights which utilized the Peacekeeper Stage 1 motor and same 92 in. fairing. The data was adjusted to account for differences in vehicle trajectories and a 3 dB uncertainty was added per MIL-STD-1540 guidelines. The resulting acoustic level is shown in Figure 4-6. 4.4.
Payload Shock Environment The maximum shock response spectrum at the base of the payload from all launch vehicle events should not exceed the flight limit levels in Figure 4-7 (clamp-band separation system). Lower separation shock levels can be achieved by the use of a Lightband separation system provided by Planetary Systems Incorporated. The resulting levels are also shown in Figure 4-7. For missions that do not utilize an Orbitalsupplied payload separation system, the shock response spectrum at the base of the payload from vehicle events should not exceed the levels in Figure 4-7 (non-separating shock).
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Figure 4-4. Minotaur IV Payload Random Vibration Environment
Figure 4-5. Minotaur IV Payload Sine Vibration Environment Release 1.1
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Figure 4-6. Minotaur IV Payload Acoustic Maximum Predicted Environment (MPE)
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Figure 4-7. Minotaur IV Payload Shock Maximum Predicted Environment (MPE) – Launch Vehicle to Payload
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If the payload employs a non-Orbital separation system, then the shock delivered from the separation systems to the Stage 4 vehicle interface must not exceed the limit level characterized in Figure 4-8. Shock above this level could require requalification of components or an acceptance of risk by the Rocket Systems Launch Program (RSLP). 4.5.
Payload Structural Integrity and Environments Verification The primary support structure for the spacecraft must possess sufficient strength, rigidity, and other characteristics required to survive the critical loading conditions that exist within the envelope of handling and mission requirements, including worst-case predicted ground, flight, and post-boost loads. It must survive those conditions in a manner that assures safety and that does not reduce the mission success probability. Spacecraft design loads are defined as follows:
a. Design Limit Load — The maximum predicted ground-based, powered flight or on-orbit load, including all uncertainties. b. Design Yield Load — The Design Limit Load multiplied by the recommended Yield Factor of Safety (YFS). The payload structure must have sufficient strength to withstand simultaneously the yield loads, applied temperature, and other accompanying environmental phenomena for each design condition without experiencing detrimental yielding or permanent deformation. c. Design Ultimate Load — The Design Limit Load multiplied by the recommended Ultimate Factor of Safety (UFS). The payload structure must have sufficient strength to withstand simultaneously the ultimate loads, applied temperature, and other accompanying environmental phenomena without experiencing any fracture or other failure mode of the structure.
Figure 4-8. Payload Shock Environment – Payload to Launch Vehicle Release 1.1
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4.5.1. Recommended Payload Testing and Analysis Sufficient payload testing and/or analysis must be performed to ensure the safety of ground crews and to ensure mission success. The payload structural design should comply with the testing and design factors of safety in Figure 4-9. Vibration testing should be based on the standard margins defined in Figure 4-10. At a minimum, it is recommended that the following tests be performed: a. Structural Integrity — Static loads, sine vibration, or other tests should be performed that combine to encompass the acceleration load environment presented in Section 4.1. b. Random Vibration — The flight level environment and recommended test level is defined in Section 4.2.1. c. Acoustics — Full scale acoustic testing is recommended to verify higher frequency dynamics of the spacecraft are not adversely affected. The acoustic levels are defined in Figure 4-6. d. Shock – The payload separation event should be simulated to verify the spacecraft is not adversely effected. Shock levels are defined in Section 4.4. The payload organization must provide OSP with a list of the tests and test levels to which the payload was subjected prior to payload arrival at the integration facility. 4.6. Thermal and Humidity Environments The thermal and humidity environment to which the payload may be exposed during vehicle processing and pad operations are defined in the sections that follow and listed in Figure 4-11. 4.6.1. Ground Operations Upon encapsulation within the fairing and for the remainder of ground operations, the payload environment will be maintained by a Heating, Ventilation and Air Conditioning (HVAC) system.
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Figure 4-9. Factors of Safety Payload Design and Test
Figure 4-10. Recommended Payload Testing Requirements
The HVAC provides conditioned air to the payload in the PPF after fairing integration. The HVAC is used at the launch pad after vehicle stacking operations. Air Conditioning (AC) is not provided during transport or lifting operations without the enhanced option that includes High Efficiency Particulate Air (HEPA) filtration. The conditioned air enters the fairing at a location forward of the payload, exits aft of the payload and is provided up to 5 minutes prior to launch. Baffles are provided at the air conditioning inlet to reduce impingement velocities on the payload if required.
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Figure 4-11. Payload Thermal and Humidity Environment Fairing inlet conditions are selected by the customer, and are bounded as follows: a. Dry Bulb Temperature: 55 to 85 °F (13 to 29 °C) controllable to ±4 °F (±2 °C) of setpoint b. Dew Point Temperature: 38 to 62 °F (3 to 17 °C) c. Relative Humidity: determined by drybulb and dewpoint temperature selections and generally controlled to within ±15%. Relative humidity is bound by the psychrometric chart and will be controlled such that the dew point within the fairing is never reached. d. Maximum Flow: 500 cfm
than 0.1. This temperature limit envelopes the maximum temperature of any component inside the payload fairing with a view factor to the payload with the exception of the Stage 4 motor. The maximum Stage 4 motor surface temperature exposed to the payload will not exceed 350 °F (177 °C), assuming no shielding between the aft end of the payload and the forward dome of the motor assembly. The Payload Adapter Module (PAM), used with the fairing to provide encapsulation of the payload during ground processing, provides some level of shielding between the payload and Stage 4 motor. Whether this temperature is attained prior to payload separation is dependent upon mission timeline.
4.6.2. Powered Flight The maximum fairing inside wall temperature will be maintained at less than 200 °F (93 °C), with an emissivity of 0.92 in the region of the payload. As a non-standard service, a low emissivity coating can be applied to reduce emissivity to less
The fairing peak vent rate is typically less than 0.6 psi/sec. Fairing deployment will be initiated at a time in flight that the maximum dynamic pressure is less than 0.01 psf or the maximum free molecular heating rate is less than 0.1 BTU/ft2/sec, as required by the payload.
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4.6.3. Nitrogen Purge (Non-Standard Service) If required for spot cooling of a payload component, Orbital will provide GN2 flow to localized regions in the fairing as a non-standard service. This option is discussed in more detail in Section 8.3.2. 4.7. Payload Contamination Control All payload integration procedures, and Orbital’s contamination control program have been designed to minimize the payload’s exposure to contamination from the time the payload arrives at the payload processing facility through orbit insertion and separation. The payload is fully encapsulated within the fairing and Payload Adapter Module (PAM) at the payload processing facility, assuring the payload environment stays clean in a Class 100,000 environment. All SLV assemblies that affect cleanliness within the encapsulated payload volume include the fairing and the payload cone assembly. These assemblies are cleaned such that there is no particulate or non-particulate matter visible to the normal unaided eye when inspected from 2 to 4 feet under 50 ft-candle incident light (Visibly Clean Level II). After encapsulation, the fairing envelope is either sealed or maintained with a positive pressure, Class 100,000 continuous purge of conditioned air. If required, the payload can be provided with enhanced contamination control as an enhanced option, providing a Class 10,000 environment, low outgassing, and Visibly Clean Plus Ultraviolet cleanliness (see Section 8.3.3). Provisions exist in the fairing design accommodate dry nitrogen purge as has been demonstrated on the Taurus application. With the enhanced contamination control option, the Orbital-supplied elements will be cleaned and controlled to support a Class 10,000 clean room environment, as defined in Federal Standard 209. This includes limiting volatile hydrocarbons to maintain hydrocarbon content at less than 15 ppm and humidity between 35 to 60
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percent. Since the payload processing will be at a GFP facility, it is assumed the Class 10,000 clean room environment also adhering to these levels of control will be provided by that facility. Also with the enhanced contamination control option, Orbital provides an Environmental Control System (ECS) from payload encapsulation through vehicle lift-off. The ECS continuously purges the fairing volume with clean filtered air. Orbital’s ECS incorporates a HEPA filter unit to provide FED-STD-209 Class M5.5 (10,000) air. Orbital monitors the supply air for particulate matter via a probe installed upstream of the fairing inlet duct prior to connecting the air source to the payload fairing. 4.8. Payload Electromagnetic Environment The payload Electromagnetic Environment (EME) results from two categories of emitters: 1) Minotaur IV onboard antennas and, 2) Range radar. All power, control and signal lines inside the payload fairing are shielded and properly terminated to minimize the potential for Electromagnetic Interference (EMI). The Minotaur IV payload fairing is Radio Frequency (RF) opaque, which shields the payload from external RF signals while the payload is encapsulated. Figure 4-12 lists the frequencies and maximum radiated signal levels from vehicle antennas that are located near the payload during ground operations and powered flight. The specific EME experienced by the payload during ground processing at the PPF and the launch site will depend somewhat on the specific facilities that are utilized as well as operational details. However, typically the field strengths experienced by the payload during ground processing with the fairing in place are controlled procedurally and will be less than 2 V/m from continuous sources and less than 10 V/m from pulse sources. The highest EME during powered flight is created by the CBand transponder transmission, which results in
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peak levels at the payload interface plane of 88 V/m at 5765 MHz (based on Taurus). Range transmitters are typically controlled to provide a field strength of 10 V/m or less inside the fairing.
This EME should be compared to the payload’s RF susceptibility levels (MIL-STD-461, RS03) to define margin.
Figure 4-12. Minotaur IV Launch Vehicle RF Emitters and Receivers
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5. PAYLOAD INTERFACES This section describes the available mechanical, electrical and Launch Support Equipment (LSE) interfaces between the Minotaur IV launch vehicle and the payload. 5.1. Payload Fairing Orbital’s flight-proven Taurus 92-inch diameter payload fairing is used to encapsulate the payload, provide protection and contamination control during ground handling, integration operations and flight. The fairing is a bi-conic design made of graphite/epoxy face sheets with aluminum honeycomb core. The two halves of the fairing are structurally joined along their longitudinal interface using Orbital’s low contamination frangible joint system. An additional circumferential frangible joint at the base of the fairing supports the fairing loads. At separation, a gas pressurization system is activated to pressurize the fairing deployment thrusters. The fairing halves then rotate about external hinges that control the fairing deployment to ensure that payload and launch vehicle clearances are maintained. All elements of the deployment system have been demonstrated through test to comply with stringent contamination requirements.
payload deflection must be given with the Finite Element Model to evaluate payload dynamic deflection with the Coupled Loads Analysis (CLA). The payload contractor should assume that the interface plane is rigid; Orbital has accounted for deflections of the interface plane. The CLA will provide final verification that the payload does not violate the dynamic envelope. 5.1.2. Payload Access Door Orbital provides one 18 in. by 24 in. (45.7 cm by 61.0 cm) payload fairing access door to provide access to the payload after fairing mate. The door can be positioned according to payload requirements within the cylindrical section of the fairing, providing access to the payload without having to remove any portion of the fairing or break electrical connections. The specific location is defined and controlled in the payload ICD. Additional access doors can readily be provided as an enhanced option (see Section 8.1.2). 5.2. Payload Mechanical Interface and Separation System Minotaur IV provides for a standard nonseparating payload interface and an optional Orbital-provided payload separation system.
5.1.1. Payload Dynamic Design Envelope The fairing drawing in Figure 5-1 shows the maximum dynamic envelope available for the payload during powered flight. The dynamic envelope shown account for fairing and vehicle structural deflections only. The payload contractor must consider deflections due to spacecraft design and manufacturing tolerance stack-up within the dynamic envelope. Proposed payload dynamic envelope violations must be approved by OSP via the ICD. No part of the payload may extend aft of the payload interface plane without specific OSP approval. Incursions below the payload interface plane may be approved on a case-by-case basis after additional verification that the incursions do not cause any detrimental effects. Vertices for
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Figure 5-1. Standard 92 in. Fairing Envelope
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Orbital will provide all flight hardware and integration services necessary to attach nonseparating and separating payloads to Minotaur IV. Payload ground handling equipment is typically the responsibility of the payload contractor. All attachment hardware, whether Orbital or customer provided, must contain locking features consisting of locking nuts, inserts or fasteners. 5.2.1. Standard Non-Separating Mechanical Interface Orbital’s payload interface design provides a standard interface that will accommodate multiple payload configurations. Figure 5-2 illustrates the standard, non-separating payload mechanical interface. This is for payloads that provide their own separation system or payloads that will not separate. The interface is a standardized circular bolted interface common with the Evolved Expendable Launch Vehicle (EELV). The interface is a 62.01-inch diameter bolted interface. A butt joint with 121 holes (0.265-inch diameter) designed for ¼-inch fasteners is the payload mounting surface as shown in Figure 5-2. Alternate or multiple payload configurations can also be accommodated with the use of a bulkhead which allows flexibility in mounting patterns and configurations. 5.2.2. Orbital Supplied Mechanical Interface Control Drawing Orbital will provide a toleranced Mechanical Interface Control Drawing (MICD) to the payload contractor to allow accurate machining of the fastener holes. The Orbital provided MICD is the only approved documentation for drilling the payload interface. 5.3. Payload Electrical Interfaces The payload electrical interface supports battery charging, external power, discrete commands, discrete telemetry, analog telemetry, serial communication, payload separation indications, and up to 16 separate ordnance discretes. If an optional Orbital-provided
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separation system is utilized, Orbital will provide all the wiring through the separable interface plane. If the option is not exercised the customer will be responsible to provide the wiring from the spacecraft to the separation plane. 5.3.1. Payload Umbilical Interfaces The payload umbilical connector provides 60 wires from the ground to the spacecraft via a dedicated payload umbilical within the vehicle, as shown in Figure 5-3. The length of the internal umbilical is approximately 25 ft (7.62 m). The cabling from the LEV to the launch vehicle is approximately 130 ft (39.6 m). This umbilical is a dedicated pass through harness for ground processing support. It allows the payload command, control, monitor, and power to be easily configured per each individual user’s requirements. The umbilical wiring is configured as a one-to-one from the Payload/Minotaur IV interface through to the payload EGSE interface in the Launch Equipment Vault, the closest location for operating customer supplied payload EGSE equipment. It is a Launch Vehicle requirement that the payload provide two (2) separation loopback circuits on the payload side of the separation plane. These are typically wired into different separation connectors for redundancy. These breakwires are used for positive separation indication telemetry and initiation of the CCAM maneuver. Figure 5-4 details the pin outs for the standard interface umbilical. All wires are twisted, shielded pairs, and pass through the entire umbilical system, both vehicle and ground, as one-to-one to simplify and standardize the payload umbilical configuration requirements while providing maximum operational flexibility to the payload provider.
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Figure 5-2. Standard, Non-separating Payload Mechanical Interface
Figure 5-3. Payload Electrical Interface Block Diagram, With No Orbital Supplied Separation System
Figure 5-4. Payload 1:1 Umbilical Pin Outs
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5.3.2. Payload Interface Circuitry Standard interface circuitry passing through the payload-to-launch vehicle electrical connections are shown in Figure 5-5. This figure details the interface characteristics for launch vehicle commands, discrete and analog telemetry, separation loopbacks, pyro initiation, and serial communications interfaces with the launch vehicle avionics systems. 5.3.3. Payload Battery Charging Orbital provides the capability for remote controlled charging of payload batteries, using a customer provided battery charger. This power is routed through the payload umbilical cable. Up to 5.0 amperes per wire pair can be accommodated. The payload battery charger should be sized to withstand the line loss from the LEV to the spacecraft. 5.3.4. Payload Command and Control The Minotaur IV standard interface provides discrete sequencing commands generated by the launch vehicle’s Ordnance Driver Module (ODM) that are available to the payload as closed circuit opto-isolator command pulses of 5 A in lengths of 35 ms minimum. The total number of ODM discretes is sixteen (16) and can be used for any combination of (redundant) ordnance events and/or discrete commands depending on the payload requirements. 5.3.5. Pyrotechnic Initiation Signals Orbital provides the capability to directly initiate 16 separate pyrotechnic conductors through two dedicated MACH Ordnance Driver Modules (ODM). Each ODM provides for up to eight drivers capable of a 5 A, 100 ms, current limited pulse into a 1.5 ohm resistive load. All eight channels can be fired simultaneously with an accuracy of 1 ms between channels. In addition, the ODM channels can be utilized to trigger high impedance discrete events if required. Safing for all payload ordnance events will be accomplished through an Arm/Disarm (A/D) Switch.
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Figure 5-5. Minotaur IV Payload Electrical Interface Block Diagram 5.3.6. Payload Telemetry The baseline telemetry subsystem capability provides a number of dedicated payload discrete (bi-level) and analog telemetry monitors through dedicated channels in the vehicle encoder. Up to 24 channels will be provided with type and data rate being defined in the mission requirements document. In addition, a GCI610 will be utilized in the encoder stack for serial data ranging up to 600 Kbs if required. The GCI610 utilizes an IRIG standard RS422 driver interface for simplicity in payload interface definition. The payload serial and analog data will be embedded in the baseline vehicle telemetry format. For discrete monitors,
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the payload customer must provide the 5 Vdc source and the return path. The current at the payload interface must be less than 10 mA. Separation breakwire monitors can be specified if required. The number of analog channels available for payload telemetry monitoring is dependent on the frequency of the data. Payload telemetry requirements and signal characteristics will be specified in the Payload ICD and should not change once the final telemetry format is released at approximately L-6 months. Orbital will tape, archive, and reduce the data into a digital format for delivery to the payloaders for review. Due to the use of strategic assets, Minotaur IV telemetry is subject to the provisions of the Strategic Arms Reduction Treaty (START). START treaty provisions require that certain Minotaur IV telemetry be unencrypted and provided to the START treaty office for dissemination to the signatories of the treaty. The extent to which START applies to the payload telemetry will be determined by RSLP. Encrypted payload telemetry can be added as a non-standard service pending approval by RSLP and the START treaty office. 5.3.7. Payload Separation Monitor Loopbacks Separation breakwire monitors are required on both sides of the payload separation plane. With the Orbital provided separation systems, Minotaur IV provides three (3) separation loopbacks on the launch vehicle side of the separation plane for positive payload separation indication.
(programmable) and data transmission bit rates. The number of channels, sample rates, etc. will be defined in the Payload ICD. 5.3.9. Non-Standard Electrical Interfaces Non-standard services such as serial command and telemetry interfaces can be negotiated between OSP and the payload contractor on a mission-by-mission basis. The selection of the separation system could also impact the payload interface design and will be defined in the Payload ICD. 5.3.10. Electrical Launch Support Equipment Orbital will provide space for a rack of customer supplied EGSE in the LCR, or either of the on-pad equipment vaults. The equipment will interface with the launch vehicle/spacecraft through either the dedicated payload umbilical interface or directly through the payload access door. The payload customer is responsible for providing cabling from the EGSE location to the launch vehicle/spacecraft. Separate payload ground processing harnesses that mate directly with the payload can be accommodated through the payload access door(s) as defined in the Payload ICD. 5.4. Payload Design Constraints The following sections provide design constraints to ensure payload compatibility with the Minotaur IV system.
Minotaur IV also requires two (2) separate loopbacks on the payload side of the separation plane. These are used for telemetry indication of separation and also the initiation of the Stage 4 CCAM maneuver.
5.4.1. Payload Center of Mass Constraints Along the Y and Z-axes, the payload CG must be within 1.0 inch (3.8 cm) of the vehicle centerline. Payloads whose CG extend beyond the 1.0 inch lateral offset limit will require Orbital to verify the specific offsets that can be accommodated.
5.3.8. Telemetry Interfaces The standard Minotaur IV payload interface provides a 16Kbps RS-422/RS-485 serial interface for payload use with the flexibility to support a variety of channel/bit rate requirements, and provide signal conditioning, PCM formatting
5.4.2. Final Mass Properties Accuracy The final mass properties statement must specify payload weight to an accuracy of at least 1 lbm (0.5 kg), the center of gravity to an accuracy of at least 0.25 inch (6.4 mm) in each axis, and the
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products of inertia to an accuracy of at least 0.5 slug-ft2 (0.7 kg-m2) (see Figure 5-6). In addition, if the payload uses liquid propellant, the slosh frequency must be provided to an accuracy of 0.2 Hz, along with a summary of the method used to determine slosh frequency.
Figure 5-6. Payload Mass Properties Measurement Tolerance 5.4.3. Pre-Launch Electrical Constraints Prior to launch, all payload electrical interface circuits are constrained to ensure there is no current flow greater than 10 mA across the payload electrical interface plane. The primary support structure of the spacecraft shall be electrically conductive to establish a single point electrical ground. 5.4.4. Payload EMI/EMC Constraints The Minotaur IV avionics share the payload area inside the fairing such that radiated emissions compatibility is paramount. OSP places no firm radiated emissions limits on the payload other than the prohibition against RF transmissions within the payload fairing. Prior to launch, Orbital requires review of the payload radiated emission levels (MIL-STD-461, RE02) to verify overall launch vehicle EMI safety margin (emission) in accordance with MIL-E-6051. Payload RF transmissions are not permitted after fairing mate and prior to an ICD specified time after separation of the payload. An EMI/EMC analysis may be required to ensure RF compatibility. Payload RF transmission frequencies must be coordinated with Orbital and range officials to ensure non-interference with Minotaur IV and range transmissions. Additionally, the customer Release 1.1
must schedule all RF tests at the integration site with Orbital in order to obtain proper range clearances 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 25 Hz (based on Taurus). However, this value is effected significantly by other factors such as incorporation of a spacecraft isolation system and/or separation system. Therefore, the final determination of compatibility must be made on a mission-specific basis. 5.4.6. Payload Propellant Slosh A slosh model should be provided to Orbital 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 RSLP. 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 for Minotaur IV payloads. These are compliance documents and must be strictly followed. It is the responsibility of the customer to ensure that the payload meets all OSP, Orbital, and range imposed safety standards. Customers designing payloads that employ hazardous subsystems are advised to contact OSP early in the design process to verify compliance with system safety standards.
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6. MISSION INTEGRATION 6.1. Mission Management Approach The Minotaur IV program is managed through U.S. Air Force, Space and Missile Systems Center, Rocket Systems Launch Program (RSLP). RSLP serves as the primary point of contact for the payload customers for the Minotaur IV launch service. A typical integrated OSP organizational structure is shown in Figure 6-1. Open communication between RSLP , Orbital, and the customer, emphasizing timely transfer of data and prudent decision-making, ensures efficient launch vehicle/payload integration operations. 6.1.1. RSLP Mission Responsibilities The program office for all OSP missions is the RSLP . They are the primary Point of Contact (POC) for all contractual and technical coordination. RSLP contracts with Orbital to provide the Launch Vehicle and launch integration and separately with commercial Spaceports and/or Government Launch Ranges for launch site facilities and services. Once a mission is identified, RSLP will assign a Mission Manager to
coordinate all mission planning and contracting activities. RSLP is supported by Northrop Grumman and other associate contractors for technical and logistical support, particularly utilizing their extensive expertise and background knowledge of the Peacekeeper booster and subsystems. 6.1.2. Orbital Mission Responsibilities As the launch vehicle provider, Orbital’s responsibilities fall into four primary areas: a. Launch Vehicle Program Management b. Mission Management c. Engineering d. Launch Site Operations Orbital assigns a Mission Manager to manage the launch vehicle technical and programmatic interfaces for a particular mission. The Orbital Mission Manager is the single POC for all aspects of a specific mission. This person has overall program authority and responsibility to ensure that payload requirements are met and that the appropriate launch vehicle services are provided. The Orbital Mission Manager will jointly chair the
Figure 6-1. Typical Integrated OSP Organizational Structure Release 1.1
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Mission Integration Working Groups (MIWGs) with the RSLP Mission Manager. The Mission Managers responsibilities include detailed mission planning, payload integration services, systems engineering, mission-peculiar design and analyses coordination, payload interface definition, launch range coordination, integrated scheduling, launch site processing, and flight operations. 6.2. Mission Planning and Development OSP will assist the customer with mission planning and development associated with Minotaur IV launch vehicle systems. These services include interface design and configuration control, development of integration processes, launch vehicle analyses, facilities planning, launch campaign planning to include range services and special operations, and integrated schedules. The procurement, analysis, integration and test activities required to place a customer’s payload into orbit are typically conducted over a 20 month standard sequence of events called the Mission Cycle. This cycle normally begins 18 months before launch, and extends to 8 weeks after launch. Once contract authority to proceed is received, the Mission Cycle is initiated. The contract option designates the payload, launch date, and basic mission parameters. In response, the Minotaur IV Program Manager designates an Orbital Mission Manager who ensures that the launch service is supplied efficiently, reliably, and on-schedule. The typical Mission Cycle interweaves the following activities: a. Mission management, document exchanges, meetings, and formal reviews required to coordinate and manage the launch service. b. Mission analyses and payload integration, document exchanges, and meetings. c. Design, review, procurement, testing and integration of all mission-peculiar hardware and software.
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d. Range interface, safety, and flight operations activities, document exchanges, meetings and reviews. Figure 6-2 details the typical Mission Cycle for a specific launch and how this cycle folds into the Orbital vehicle production schedule with typical payload activities and milestones. A typical Mission Cycle is based on an 18 month interval between mission authorization and launch. This interval reflects the OSP contractual schedule and has been shown to be an efficient schedule based on Orbital’s Minotaur, Taurus and Pegasus program experience. However, OSP is flexible to negotiate either accelerated cycles, which take advantage of the Minotaur IV/Pegasus/Minotaur/ Taurus multi-customer production sets, or extended cycles required by unusual payload requirements, such as extensive analysis or complex payload-launch vehicle integrated designs or tests or funding limitations. 6.3. Mission Integration Process 6.3.1. Integration Meetings The core of the mission integration process consists of a series of Mission Integration and Range Working Groups (MIWG and RWG, respectively). The MIWG has responsibility for all physical interfaces between the payload and the launch vehicle. As such, the MIWG creates and implements the Payload-to-Minotaur IV ICD in addition to all mission-unique analyses, hardware, software, and integrated procedures. The RWG is responsible for the areas of launch site operations; range interfaces; safety review and approval; and flight design, trajectory, and guidance. Documentation produced by the RWG includes all required range and safety submittals. Working Group membership consists of the Mission Manager and representatives from Minotaur IV engineering and operations organizations, as well as their counterparts from the customer organization. While the number of meetings, both formal and informal, required to develop and implement the mission integration
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Figure 6-2. Typical Mission Integration Schedule process will vary with the complexity of the spacecraft, quarterly meetings are typical. 6.3.2. Mission Design Reviews (MDR) Two mission-specific design reviews will be held to determine the status and adequacy of the launch vehicle mission preparations. They are designated MDR-1 and MDR-2 and are typically held 6 months and 13 months, respectively, after authority to proceed. They are each analogous to Preliminary Design Reviews (PDRs) and Critical Design Reviews (CDRs), but focus primarily on mission-specific elements of the launch vehicle effort. 6.3.3. Readiness Reviews During the integration process, reviews are held to provide the coordination of mission participants and management outside of the regular contact of the Working Groups. Due to the
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variability in complexity of different payloads and missions, the content and number of these reviews can be tailored to customer requirements. As a baseline, Orbital will conduct two readiness reviews as described below. a. Mission Readiness Review — Conducted within 1 month of launch, the Mission Readiness Review (MRR) provides a pre-launch assessment of integrated launch vehicle/payload/facility readiness prior to committing significant resources to the launch campaign. b. Launch Readiness Review — The Launch Readiness Review (LRR) is conducted at L-1 day and serves as the final assessment of mission readiness prior to activation of range resources on the day of launch.
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6.4. Documentation Integration of the payload requires detailed, complete, and timely preparation and submittal of interface documentation. As the launch service provider, RSLP is the primary communication path with support agencies, which include—but are not limited to—the various Range support agencies and U.S. Government agencies such as the U.S. Department of Transportation and U.S. State Department. Customer-provided documents represent the formal communication of requirements, safety data, system descriptions, and mission operations planning. The major products and submittal times associated with these organizations are divided into two areas— those products that are provided by the customer, and those produced by Orbital. 6.4.1. Customer-Provided Documentation Documentation produced by the customer is detailed in the following paragraphs. 6.4.1.1. Payload Questionnaire The Payload Questionnaire is designed to provide the initial definition of payload requirements, interface details, launch site facilities, and preliminary safety data to OSP. The customer shall provide a response to the Payload Questionnaire form (Appendix A), or provide the same information in a different format, in time to support the Mission Kickoff Meeting. The customer’s responses to the payload questionnaire define the most current payload requirements and interfaces and are instrumental in Orbital’s preparation of numerous documents including the ICD, Preliminary Mission Analysis, and launch range documentation. Additional pertinent information, as well as preliminary payload drawings, should also be included with the response. Orbital understands that a definitive response to some questions may not be feasible. These items are defined during the normal mission integration process. 6.4.1.2. Payload Mass Properties Payload mass properties must be provided in a timely manner in order to support efficient launch vehicle trajectory development and dynamic Release 1.1
analyses. Preliminary mass properties should be submitted as part of the MRD at launch vehicle authority to proceed. Updated mass properties shall be provided at predefined intervals identified during the initial mission integration process. Typical timing of these deliveries is included in Figure 6-2. 6.4.1.3. Payload Finite Element Model A payload mathematical model is required for use in Orbital’s preliminary coupled loads analyses. Acceptable forms include either a Craig-Bampton model valid to 120 Hz or a NASTRAN finite element model. For the final coupled loads analysis, a test verified mathematical model is desired. 6.4.1.4. Payload Thermal Model for Integrated Thermal Analysis An integrated thermal analysis can be performed for any payload as a non-standard service. A payload thermal model will be required from the payload organization for use in Orbital’s integrated thermal analysis if it is required. The analysis is conducted for three mission phases: a. Prelaunch ground operations; b. Ascent from lift-off until fairing jettison; and c. Fairing jettison through payload deployment. Models must be provided in SINDA format. There is no limit on model size although turnaround time may be increased for large models. 6.4.1.5. Payload Drawings Orbital prefers electronic versions of payload configuration drawings to be used in the mission specific interface control drawing, if possible. Orbital will work with the customer to define the content and desired format for the drawings. 6.4.1.6. Program Requirements Document (PRD) Mission Specific Annex Inputs To obtain range support, a PRD must be prepared. This document describes requirements needed to generally support the Minotaur IV launch vehicle. For each launch, an annex is submitted to specify the range support needed to
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meet the mission’s requirements. This annex includes all payload requirements as well as any additional Minotaur IV requirements that may arise to support a particular mission. The customer completes all appropriate PRD forms for submittal to Orbital. 6.4.1.6.1. Launch Operations Requirements (OR) Inputs To obtain range support for the launch operation and associated rehearsals, an OR must be prepared. The customer must provide all payload pre-launch and launch day requirements for incorporation into the mission OR. 6.5. Safety 6.5.1. System Safety Requirements In the initial phases of the mission integration effort, regulations and instructions that apply to spacecraft design and processing are reviewed. Not all safety regulations will apply to a particular mission integration activity. Tailoring the range requirements to the mission unique activities will be the first step in establishing the safety plan. OSP has three distinctly different mission approaches affecting the establishment of the safety requirements: a. Baseline mission: Payload integration and launch operations are conducted at VAFB, CA b. Campaign/VAFB Payload Integration mission: Payload integration is conducted at VAFB and launch operations are conducted from a non-VAFB launch location. c. Campaign/Non-VAFB Payload Integration mission: Payload integration and launch operations are conducted at a site other than VAFB. For the baseline and VAFB Payload Integration missions, spacecraft prelaunch operations are conducted at Government Furnished Property (GFP) Payload Processing Facility (PPF). For campaign style missions, the spacecraft prelaunch operations are performed at the desired launch site. Release 1.1
Before a spacecraft arrives at the processing 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 Requirements for baseline and VAFB Payload Integration 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. Processing and launch site ranges discourage the use of waivers and variances. Furthermore, approval of such waivers cannot be guaranteed. 6.5.2. System Safety Documentation For each Minotaur IV 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 respective systems should be included. Major categories of hazardous systems are ordnance devices, radioactive materials, propellants, pressurized systems, toxic materials, cryogenics, and RF radiation. Procedures relating to these systems as well as any procedures relating to lifting operations or battery operations should be prepared for safety review submittal. OSP will provide this information to the appropriate safety offices for approval.
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7. GROUND AND LAUNCH OPERATIONS 7.1. Minotaur IV/Payload Integration Overview The processing of the Minotaur IV utilizes many of the same proven techniques developed for the Pegasus, Taurus and Minotaur launch vehicles. This minimizes the handling complexity for both vehicle and payload. 7.2. Ground And Launch Operations Ground and launch operations are conducted in three major phases: a. Launch Vehicle Integration — Assembly and test of the Minotaur IV vehicle b. Payload Processing/Integration — Receipt and checkout of the satellite payload, followed by integration with Minotaur IV fairing and Payload Adapter Module (PAM) and verification of interfaces c. Launch Operations — Includes transport to the launch pad, final integration, checkout, arming and launch. 7.2.1. Launch Vehicle Integration Orbital will process all Minotaur IV vehicles according to a flow similar to that implemented for the Minotaur and Taurus vehicles. All launch vehicle motors, parts and completed subassemblies are delivered to the launch Vehicle Processing Facility (VPF) from Orbital’s Chandler production facility, the assembly/motor vendor or the Government. Figure 7-1 depicts the typical flow of hardware from the factory to the launch site. Flowcharts of the field processing are shown in Figure 7-2. 7.2.1.1. Planning and Documentation Minotaur IV integration and test activities are controlled by a comprehensive set of Work Packages (WPs) that describe and document every aspect of integrating and testing Minotaur IV and its payload. All testing and integration activities are scheduled by work package number on a daily activity schedule updated and distributed daily during field operations. This schedule is maintained by Orbital and serves as Release 1.1
the master document communicating all activities planned at the field site. The schedule contains notations regarding the status of the work package document and hardware required to begin the operation. Mission-specific work packages are created for mission-unique or payload-specific procedures. Any discrepancies encountered are recorded on a Discrepancy Report and dispositioned as required. All activities are in accordance with Orbital’s ISO 9001 certification. 7.2.1.2. GCA/Orion 38 Integration and Test Activities The GCA will undergo subsystem level testing at Orbital’s Chandler facility prior to being shipped to the field site. The GCA and the Stage 4 Orion 38 motor are then delivered to the launch vehicle processing facility located at VAFB. Upon arrival at VAFB these components/sub-assemblies will undergo a thorough inspection and subsystem level checkout. At this time range certification of Range Tracking System (RTS) and Flight Termination System (FTS) devices will be performed. The components will be reinstalled and in-vehicle testing of the RTS and FTS systems will be performed. After the completion of subsystem level testing the Orion 38 motor is integrated into the GCA to form the Stage 4 assembly. 7.2.1.3. PK Motor Integration and Test Activities The PK Stage 1, 2 and 3 motors are delivered to the launch vehicle processing facility where they undergo checkout and testing. Once integration is complete, a booster confidence test will be conducted. 7.2.1.4. Mission Simulation Tests Orbital will run three Mission Simulator Tests (MST) to verify the functionality of launch vehicle hardware, and software. The Mission Simulation Tests use the actual flight software and simulate a “fly to orbit” scenario using simulated Inertial Navigation System (INS) data. This will allow the test to proceed throughout all mission phases
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Figure 7-1. Hardware Flow – Factory to Launch Site
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Figure 7-2. SLV Processing Flow recording vehicle performance data. The data will be compared to simulations performed in the factory software laboratory using an identical copy of the flight software. Orbital will use GFP PK nozzle assembly simulators to perform all mission simulations. These components will provide a realistic assessment of booster performance during the testing operations. After a thorough data review of all telemetry parameters, the test configuration is disassembled and setup for payload integration begins. 7.2.1.5. Booster Assembly Stacking/Launch Pad Preparation After completion of the MST, the booster assembly (Stages 1, 2 and 3) and the stage 4 assembly (Orion 38 integrated with the GCA) are transported to the launch facility. Figure 7-3 shows a pictorial representation of the processing flow. Prior to the arrival of the PK boosters, the site is prepared for launch operations with the Release 1.1
installation of the launch stool. This stool is the same design used for Orbital’s Taurus SLV. It supports a flat pad launch of a full PK booster assembly with front section and fourth stage motor options. After stool installation, the fixed scaffolding installation is performed. This scaffolding provides access to the base ring of the PK Stage 1 motor during integration activities. Once the booster arrives at the launch site, it is then lifted and emplaced onto the launch stool. Each motor assembly will be individually stacked using a process developed for handling Taurus Stage 0 motors. Scaffolding integration continues as the booster stages are mated. The Stage 4 assembly is shipped in the vertical configuration to the launch facility for payload integration. 7.2.2. Payload Processing/Integration Payloads normally undergo initial checkout and preparation for launch at an Air Force payload
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Figure 7-3. Minotaur IV Processing Flow processing facility (PPF) or commercial facilities at VAFB. After arrival at the PPF, the payload completes its own independent verification and checkout prior to beginning integrated processing with Minotaur IV fairing and Payload Adapter Module (PAM). The Minotaur IV fairing and PAM will be delivered to the payload processing facility for encapsulation of the payload. The fairing and PAM provide a sealed enclosure which protects the payload and provides a structure to facilitate transportation to the launch facility. After enclosure of the payload in the fairing, the assembly is shipped in the vertical configuration to the launch facility for a pre-installation verification test. 7.2.2.1. Payload Propellant Loading Payloads utilizing integral propulsion systems with propellants such as hydrazine can be loaded and secured through coordinated Orbital and contractor arrangements for use of the propellant Release 1.1
loading facilities in the VAB. standard service.
This is a non-
7.2.2.2. Final Vehicle Integration and Test After successful completion of payload mate/fairing closeout the completed front section assembly (Minotaur IV Stage 4 assembly integrated with the payload assembly) will then be lifted in vertical configuration atop the booster assembly. Figure 7-4 illustrates the vertical lifting operation performed on a Taurus front section. Final post mates checks of the booster assembly and front section assembly interface are then conducted. A final systems verification test, similar to the previous MST, is then performed. At this point the vehicle is ready for final Range interface tests and launch readiness.
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Figure 7-4. Minotaur IV Upper Stack Assembly will be Vertically Integrated to Minotaur IV Booster Assembly in a Similar Manner to Taurus Upper Stack
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7.3.1. Launch Control Organization The Launch Control Organization is split into two groups: the Management group and the Technical group. The Management group consists of senior range personnel and Mission Directors/Managers for the launch vehicle and payload. The Technical Group consists of the personnel responsible for the execution of the launch operation and data review/assessment for the Payload, the Launch Vehicle and the Range. The Payload’s members of the technical group are engineers who provide technical representation in the control center. The Launch Vehicle’s members of the technical group are engineers who prepare the Minotaur IV for flight, review and assess data that is displayed in the Launch Control Room (LCR) and provide technical representation in the LCR and in the Launch Operations Control Center (LOCC). The Range’s members of the technical group are personnel that maintain 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.
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8. OPTIONAL ENHANCED CAPABILITIES The OSP launch service is structured to provide a baseline vehicle configuration which is then augmented with optional enhancements to meet the unique needs of individual payloads. The baseline vehicle capabilities are defined in the previous sections and the optional enhanced capabilities are defined below. The enhanced options allow customization of launch support and accommodations to the Minotaur IV designs on an efficient, “as needed” basis. 8.1. Mechanical Interface and Separation System Enhancements 8.1.1. Separation Systems Various separation systems can be provided or accommodated to meet mission-unique requirements. As a baseline option, Orbital offers an optional payload separation system that is flight proven on Taurus. The separation system is manufactured for Orbital by SAAB Ericson Space (SES). SES has extensive experience in supplying separation systems for a wide range of launch vehicles and payloads. This system is based on a design that has flown over 30 times with 100% success. The baseline separation system, shown in Figure 8-1, has a standardized 38.81 inch bolt pattern. It is a marmon clamp design employing two aluminum interface rings that are clamped by dual, semi-circular stainless steel clamp bands with aluminum clamp shoes. Each of the two retention bolts is severed by a redundantly initiated bolt cutter. The separation ring to which the payload attaches is supplied with through holes and the separation system is mated to the spacecraft during processing at the PPF. The weight of hardware separated with the payload is approximately 8.7 lbm (13.95 kg). Orbitalprovided attachment bolts to this interface can be inserted from either the launch vehicle or the payload side of the interface (NAS630xU, dash number based on payload flange thickness). The weight of the bolts, nuts, and washers connecting
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the separation system to the payload is allocated to the separation system and included in the launch vehicle mass. Separation velocity is provided by up to eight matched spring actuators. The spring assemblies may be tailored to mitigate the effects of payload CG offset, controlling tip-off within 5 deg/sec. Tipoff rates are highly dependent on payload mass properties, but are typically on the order 1 deg/sec. Preliminary and final mission-specific tipoff analyses are conducted for each payload using Orbital’s computer simulation dynamic analysis tools. If non-standard separation velocities are needed, different springs may be substituted on a mission-specific basis as a non-standard service. Other separation systems can supplied on a mission-specific basis.
also
be
8.1.2. Additional Fairing Access Doors Additional access doors can be provided to accommodate unique payload requirements. The standard door size is 18 in. by 24 in. (45.7 cm by 61.0 cm). Access doors of non-standard size can also be provided as necessary. Orbital performs structural analyses to verify the acceptability of the mission-specific door configuration. 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. The 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 and six Taurus missions. The typical Minotaur configuration is shown in Figure 8-2. 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 overall payload center of gravity loads and component or subsystem
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Figure 8-1. 38-in. Separation System Option responses. Typically the system will reduce transient loads to approximately 50% of the level they would be without the system. The exact results can be expected to vary for each particular spacecraft and with location on the spacecraft. Generally, a beneficial reduction in shock and vibration will also be provided. The isolation system does impact overall vehicle performance (by approximately 20 to 40 lb [9 to 18 kg]) and the available payload dynamic envelope by up to 4 inches (10.16 cm) axially and up to 1.0 inch (2.54 cm) laterally.
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8.2. Performance Enhancements 8.2.1. Insertion Accuracy Enhanced insertion accuracy or support for multiple payload insertion can be provided as an enhanced option utilizing the Hydrazine Auxiliary Propulsion System (HAPS). The common usage of the Orion 38 makes the flight proven HAPS design directly applicable to the Minotaur IV usage, as shown in Figure 8-3. Orbital insertion accuracy can typically be improved to ±10 nmi (±18.5 km) or better in each apse and ±0.05 deg in inclination. For orbits above 324 nmi (600 km), the
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Figure 8-2. Soft Ride Payload Isolation System as Integrated on Minotaur LV
HAPS also permits injection of shared payloads into different orbits. HAPS, which is mounted inside the Avionics Structure, consists of a hydrazine propulsion subsystem and a Stage 4 separation subsystem. After burnout and separation from the Stage 4 motor, the HAPS hydrazine thrusters provide additional velocity and both improved performance and precise orbit injection. The HAPS propulsion subsystem consists of a centrally mounted tank containing approximately 130 lbm (59 kg) of hydrazine, helium pressurization gas, and three fixed, axially pointed thrusters. The hydrazine tank contains an integral bladder which will support multiple restarts.
HAPS can also increase payload mass by approximately 50 to 250 lbm (22.7 to 113 kg), depending on the orbit. Specific performance capability associated with the HAPS can be provided by contacting the OSP program office.
8.2.2. Star 48 Stage 4 The modular design of Orbital’s GCA and integrating structures provides great flexibility in accommodating alternative Stage 4 propulsion systems. As one low risk example, a optional
Figure 8-3. Hydrazine Auxiliary Propulsion System (HAPS) Used to Provide Insertion Accuracy
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configuration using the more powerful ATK Thiokol Star-48 motor is shown in Figure 8-4. The only modifications required to accommodate this change are a modified Motor Adapter Cone (MAC) with the Star 48 forward interface and a longer 3/4 interstage to allow room for the increased motor length. Other alternative motors can be similarly adopted.
100,000 or Class 10,000). Nitrogen purge is used in conjunction with the Conditioned Air option to provide mission-specific localized cooling and/or dry nitrogen environments to satisfy unique payload environmental requirements. 8.3.2. Nitrogen Purge Continuous clean dry nitrogen inside the shroud during vehicle processing from payload encapsulation to launch is available as an option. Dry clean nitrogen purge can be provided to the payload at a Class 10,000 environment for continuous purge of the payload after fairing encapsulation until lift-off. The capability was demonstrated on the Minotaur MightySat mission with the exception of purge during transportation. A nitrogen cooling system is already provided on every SLV mission to spot-cool sensitive electronic boxes. Flow adjustments for cooling versus purge would be changed back and forth to accommodate both. If nitrogen purge is required during transport, the only additional items needed would be a minor addition of a nitrogen bottle, regulator, and mounting hardware.
Figure 8-4. Orion 38 Stage 4 Motor can be Replaced with a Star-48 to Provide Increased Performance 8.3. Environmental Control Options 8.3.1. Conditioned Air Conditioned air can be provided within the fairing volume using an Environmental Control System (ECS) via a “fly-out” duct that is retracted at launch. Temperature and humidity is regulated within the limits specified in the Payload ICD. A filter is installed to provide a Class 100,000 environment, typically. The Nitrogen Purge (Section 4.6.3) and Enhanced Contamination Control (Section 4.7 and 8.3.3) enhancements complement this capability. Upon exercise of the Enhanced Cleanliness option, a certified HEPA filter is used in the input duct to assure the necessary low particulate environment (Class Release 1.1
The system distribution lines are routed across the payload interface plane and/or along the inner surface of the shroud or fairing. If required for spot cooling of a payload component, Orbital will provide GN2 flow to localized regions in the fairing. The GN2 will meet Grade B specifications, as defined in MIL-P-27401C and can be regulated to at least 5 scfm. The system’s regulators are set to a desired flow rate during pre-launch processing. Payload purge requirements are controlled and documented via the launch vehicle to payload ICD. Payload purge requirements must be coordinated with Orbital via the ICD to ensure that the requirements can be achieved. 8.3.3. Enhanced Contamination Control Understanding that some payloads have requirements for enhanced cleanliness, OSP offers a contamination control option, which is
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composed of the elements in the following sections (which is also discussed in Section 4.7). Minotaur IV customers can also coordinate combinations of the elements listed below to meet the unique needs of their payloads. 8.3.3.1. High Cleanliness Integration Environment (Class 10K or 100K) With enhanced contamination control, a soft walled clean room can be provided to ensure a FED-STD-209 Class M6.5 (100,000) or Class M5.5 (10,000) environment during all payload processing activities up to fairing encapsulation. The soft walled clean room and anteroom(s) utilize HEPA filter units to filter the air and hydrocarbon content is maintained at 15 ppm or less. The payload organization is responsible for providing the necessary clean room garments for payload staff as well as vehicle staff that need to work inside the clean room. 8.3.3.2. Fairing Surface Cleanliness Options The inner surface of the fairing and payload cone assemblies can be cleaned to cleanliness criteria which ensures no particulate matter visible with normal vision when inspected from 6 to 18 inches under 100 ft-candle incident light. The same will be true when the surface is illuminated using black light, 3200 to 3800 Angstroms (Visibly Clean Plus Ultraviolet). In addition, Orbital can ensure that all materials used within the encapsulated volume have outgassing characteristics of less than 1.0% TML and less than 0.1% CVCM. Items that do not meet these levels can be masked to ensure they are encapsulated and will have no significant effect on the payload. 8.3.3.3. High Cleanliness Fairing Environment With the enhanced contamination control option, Orbital provides an ECS from payload encapsulation until just prior to vehicle lift-off. The ECS continuously purges the fairing volume with clean filtered air. Orbital’s ECS incorporates a HEPA filter unit to provide FED-STD-209 Class M5.5 (10,000) air. Orbital monitors the supply air for particulate matter via a probe installed Release 1.1
upstream of the fairing inlet duct prior to connecting the air source to the payload fairing. 8.3.4. Launch Pad Environmental Control For launch sites without gantries for environmental control or vehicle access, optional Quick Erect Scaffold® (QES) will protect the vehicle from the environments, maintain temperatures within 60 to 100 °F (in conjunction with a thermal blanket) and provide access to the vehicle for launch pad operations (see Figure 8-5). As the name implies, QES can be rapidly assembled and is highly adaptable for accommodating different vehicle configurations. This scaffolding has been previously demonstrated on Minotaur. 8.3.4.1. Booster Temperature Control The thermal blanket design successfully used during Minotaur missions can be used to maintain the PK booster operating temperature within the limits of 60 to 100 °F. The thermal blanket is
Figure 8-5. Mobile Scaffolding for Environmental Control Demonstrated on Minotaur Missions
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constructed of outer PVC material with an inner insulating liner. It is a four piece cover with Velcro seams running along the length of the boosters. Integral inflatable manifold tubes space the blanket away from the booster and provide space for ducting conditioned air for the boosters. The baseline blanket design only covers the Stage 1, 2 and 3 boosters. 8.4. Enhanced Telemetry Options OSP can provide mission specific instrumentation and telemetry components to support additional payload or experiment data acquisition requirements. Telemetry options include additional payload-dedicated bandwidth and GPS-based precision navigation data. 8.4.1. Enhanced Telemetry Bandwidth Enhanced mission specific instrumentation and telemetry can be provided, supplying a dedicated telemetry link to support additional payload or experiment data acquisition requirements. A baseline data rate of 1 Mbps is available, however, maximum data rates depend on the mission coverage required and the launch range receiver characteristics and configuration. The enhanced telemetry option was demonstrated on both inaugural Minotaur TLV and SLV missions. 8.4.2. Enhanced Telemetry Instrumentation To support the higher data rate capability in Section 8.4.1, enhanced telemetry instrumentation can be provided. The instrumentation can include strain gauges, temperature sensors, accelerometers, analog data, and digital data configured to mission-specific requirements. This capability was successfully demonstrated on the inaugural OSP-SLV mission. 8.4.3. Navigation Data Precision navigation data using an independent Global Positioning System (GPS) receiver and telemetry link is available as an enhanced option. This option utilizes Orbital’s flight proven GPS Position Beacon (GPB) to provide missile state data for range safety and
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provides better than 100 m position accuracy with 10 Hz data rate. This capability was successfully demonstrated on the inaugural Minotaur mission. 8.5. Shared Launch Accommodations Minotaur IV is uniquely capable of providing launches of multiple satellite payloads, leveraging RSLP and Orbital’s extensive experience in integrating and launching multiple payloads. Multiple spacecraft configurations have been flown on many of Orbital’s Pegasus, Taurus and Minotaur missions to date. A number of different structural configurations have been developed for dual payloads, one is shown in Figure 8-6. Because of the modular nature of the structures, dual payload configurations can be easily accommodated by the Minotaur IV structural design.
Figure 8-6. Modular Minotaur IV Structural Design Easily Accommodates Multiple Payloads
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Appendix A
APPENDIX A PAYLOAD QUESTIONNAIRE
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SATELLITE IDENTIFICATION FULL NAME: ACRONYM: OWNER/OPERATOR: INTEGRATOR(s): ORBIT INSERTION REQUIREMENTS* SPHEROID
Standard (WGS-84, Re = 6378.137 km) Other:
ALTITUDE Insertion Apse: nmi ___ or... Semi-Major Axis:
Opposite Apse: ±
__
km
nmi km
± Eccentricity:
nmi ___ INCLINATION
±
__
≤e≤
km
± deg
ORIENTATION Argument of Perigee:
Longitude of Ascending Node (LAN): ±
deg
±
deg
Right Ascension of Ascending Node (RAAN): ±
deg
...for Launch Date: * Note: Mean orbital elements
LAUNCH WINDOW REQUIREMENTS NOMINAL LAUNCH DATE: OTHER CONSTRAINTS (if not already implicit from LAN or RAAN requirements, e.g., solar beta angle, eclipse time constraints, early on-orbit ops, etc):
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Appendix A
GROUND SUPPORT EQUIPMENT Describe any additional control facilities (other than the baseline Support Equipment Building (SEB) and Launch Equipment Vault (LEV)) which the satellite intends to use:
SEB
Describe (in the table below) Satellite EGSE to be located in the LSV. [Note: Space limitations exist in the SEB, 350 ft umbilical cable length to spacecraft typical] Equipment Name / Type
Approximate Size (LxWxH)
Power Requirements
..................................................................
...............................................
................................
..................................................................
...............................................
................................
..................................................................
...............................................
................................
..................................................................
...............................................
................................
..................................................................
...............................................
................................
Is UPS required for equipment in the SEB? Is Phone/Fax connection required in the SEB? LEV
Yes / No Yes / No
Circle: Phone / FAX
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] Approximate Size (LxWxH) Power Requirements Equipment Name / Type ..................................................................
...............................................
................................
..................................................................
...............................................
................................
..................................................................
...............................................
................................
..................................................................
...............................................
................................
..................................................................
...............................................
................................
Is UPS required for equipment in the LEV? Is Ethernet connection between SEB and LEV required?
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EARLY ON-ORBIT OPERATIONS Briefly describe the satellite early on-orbit operations, e.g., event triggers (separation sense, sun acquisition, etc), array deployment(s), spin ups/downs, etc:
SATELLITE SEPARATION REQUIREMENTS ACCELERATION Longitudinal:
=
g’s
Lateral:
=
g’s
VELOCITY Relative Separation Velocity Constraints: ANGULAR RATES Longitudinal: (pre-separation) ±
deg/sec
ANGULAR RATES Longitudinal: (post-separation) ±
deg/sec
Pitch:
±
deg/sec
Yaw:
±
deg/sec
Pitch:
±
deg/sec
Yaw:
±
deg/sec
ATTITUDE Describe Pointing Requirements Including Tolerances: (at deployment)
SPIN UP Longitudinal Spin Rate: OTHER
±
deg/sec
Describe Any Other Separation Requirements:
SPACECRAFT COORDINATE SYSTEM Describe the Origin and Orientation of the spacecraft reference coordinate system, including its orientation with respect to the launch vehicle (provide illustration if available):
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SPACECRAFT PHYSICAL DIMENSIONS STOWED CONFIGURATION
Length/Height:
Diameter: in
cm
in
cm
Other Pertinent Dimension(s):
Describe any appendages/antennas/etc which extend beyond the basic satellite envelope:
ON-ORBIT Describe size and shape: CONFIGURATION
If available, provide dimensioned drawings for both stowed and on-orbit configurations.
SPACECRAFT MASS PROPERTIES* PRE-SEPARATION Mass: Xcg: Ycg: Zcg: POST-SEPARATION Mass: (non-separating adapter remaining with launch vehicle) Xcg: Ycg: Zcg:
Inertia units: lbm kg Ixx: in cm Izz: in cm Iyz: in cm Inertia units: lbm kg Ixx: in cm Izz: in cm Iyz: in cm
lbm-in2
kg-m2
Iyy: Ixy: Ixz: lbm-in2
kg-m2
Iyy: Ixy: Ixz:
* Stowed configuration, spacecraft coordinate frame
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Appendix A
ASCENT TRAJECTORY REQUIREMENTS Free Molecular Heating at Fairing Separation: (Standard Service: = 360 Btu/ft2/hr)
Btu/ft2/hr FMH =
W/m2
Fairing Internal Wall Temperature (Standard Service: = 200°F)
deg F T =
deg C
Dynamic Pressure at Fairing Separation: (Standard Service: = 0.01 lbf /ft2)
lbf /ft2 q =
N/m2
Ambient Pressure at Fairing Separation: (Standard Service: = 0.3 psia)
lbf /in2 P =
N/m2
Maximum Pressure Decay During Ascent: (Standard Service: = 0.6 psia)
lbf /in2/sec ΔP =
N/m2/sec
Thermal Maneuvers During Coast Periods: (Standard Service: none) SPACECRAFT ENVIRONMENTS THERMAL Spacecraft Thermal Dissipation, Pre-Launch Encapsulated: DISSIPATION Approximate Location of Heat Source: TEMPERATURE Temperature Limits During Ground/Launch Operations:
Watts
Max Min (Standard Service is 55°F to 80°F)
deg F deg F
deg C deg C
Component(s) Driving Temperature Constraint: Approximate Location(s): HUMIDITY Relative Humidity: Max % % Min
Dew Point: Max deg F Min deg F (Standard Service is 37 deg F)
or,
deg C deg C
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:
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Lateral =
g’s
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ELECTRICAL INTERFACE Bonding Requirements: Are Launch Vehicle Supplied Pyro Commands Required?
Yes / No If Yes, magnitude: (Standard Service is 10 amps for 100 msec)
Are Launch Vehicle Supplied Discrete Commands Required? Yes / No
amps for
msec
If Yes, describe:
Is Electrical Access to the Satellite Required... After Encapsulation? at Launch Site Yes / No
Yes / No
Is Satellite Battery Charging Required... After Encapsulation? at Launch Site? Yes / No
Yes / No
Is a Telemetry Interface with the Launch Vehicle Flight Computer Required?
Yes / No
If Yes, describe:
Other Electrical Requirements:
Please complete attached sheet of required pass-through signals. RF RADIATION Time After Separation Until RF Devices Are Activated: (Note: Typically, no spacecraft radiation is allowed from encapsulation until 30 minutes after liftoff.) Frequency:
MHz
Power:
Watts
Location(s) on Satellite (spacecraft coordinate frame): Longitudinal
in
cm
Clocking (deg), Describe:
Longitudinal
in
cm
Clocking (deg), Describe:
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REQUIRED PASS-THROUGH SIGNALS Item #
Pin
Signal Name
From LEV
To Satellite
Shielding
Max Current (amps)
Total Line Resistance (ohms)
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28
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Appendix A 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
Flatness Requirements for Sep System or Mating Surface of Launch Vehicle:
FAIRING Payload Fairing Access Doors (spacecraft coordinate frame): ACCESS Longitudinal in cm Clocking (deg), Describe: Longitudinal
in
cm
Clocking (deg), Describe:
Note: Standard Service is one door DYNAMICS Spacecraft Natural Frequency: Axial Recommended: OTHER
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Hz
Lateral
> TBD Hz
Hz > TBD Hz
Other Mechanical Interface Requirements:
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