Pegasus

August 2000Release 5.0

Approved for Public ReleaseDistribution Unlimited

© 2000 by Orbital Sciences Corporation. All rights reserved.G152.00

August 2000

Release 5.0

Pegasus® Users Guide

Approved for Public ReleaseDistribution UnlimitedCopyright © 1996—2000 by Orbital SciencesCorporation. All Rights Reserved.

ORBITAL SCIENCES CORPORATION

Pegasus User's Guide Preface

Release 5.0 August 2000

This Pegasus User's Guide is intended to familiarize potential space launch vehicle users with thePegasus launch system, its capabilities and its associated services. The launch services describedherein are available for commercial procurement directly from Orbital Sciences Corporation.

Readers desiring further information on Pegasus should contact us via:

E-mail to: [email protected]

Telephone: (703) 404-7400Facsimile: (703) 404-8042

Copies of this Pegasus User's Guide may be obtained from our website at http://www.orbital.com.Hardcopy documents and electronic (CD format) are also available upon request.


Pegasus User's Guide

1.0 INTRODUCTION ............................................................................................ 1-1

2.0 PEGASUS XL VEHICLE DESCRIPTION ........................................................................ 2-12.1 Pegasus XL Vehicle Description ....................................................................... 2-1

2.1.1 Solid Rocket Motors .............................................................................. 2-12.1.2 Payload Fairing ..................................................................................... 2-12.1.3 Avionics ............................................................................................ 2-22.1.4 Flight Termination System ..................................................................... 2-22.1.5 Attitude Control Systems ....................................................................... 2-32.1.6 Telemetry Subsystem............................................................................. 2-42.1.7 Major Structural Subsystems ................................................................. 2-4

2.1.7.1 Wing.................................................................................... 2-42.1.7.2 Aft Skirt Assembly ................................................................ 2-42.1.7.3. Payload Interface Systems .................................................... 2-5

2.2 Orbital Carrier Aircraft ...................................................................................... 2-5

3.0 GENERAL PERFORMANCE CAPABILITY ..................................................................... 3-13.1 Mission Profiles ............................................................................................ 3-13.2 Performance Capability .................................................................................... 3-23.3 Trajectory Design Optimization ........................................................................ 3-33.4 Injection Accuracy ............................................................................................ 3-3

3.4.1 Actual Pegasus Injection Accuracies ..................................................... 3-43.4.2 Error-Minimizing Guidance Strategies ................................................... 3-4

3.5 Collision/Contamination Avoidance Maneuver ................................................. 3-5

4.0 PAYLOAD ENVIRONMENT......................................................................................... 4-14.1 Design Loads ............................................................................................ 4-14.2 Payload Testing and Analysis ............................................................................. 4-14.3 Payload Acceleration Environment .................................................................... 4-1

4.3.1 Drop Transient Acceleration .................................................................. 4-24.4 Payload Vibration Environment ......................................................................... 4-2

4.4.1 Long Duration Captive Carry ................................................................ 4-24.5 Payload Shock Environment .............................................................................. 4-24.6 Payload Acoustic Environment .......................................................................... 4-24.7 Payload Thermal and Humidity Environment .................................................... 4-3

4.7.1 Nitrogen Purge ..................................................................................... 4-44.8 Payload Electromagnetic Environment .............................................................. 4-54.9 Payload Contamination Control ........................................................................ 4-54.10 Payload Deployment ........................................................................................ 4-94.11 Payload Tip-off .......................................................................................... 4-10

5.0 SPACECRAFT INTERFACES .......................................................................................... 5-15.1 Payload Fairing ............................................................................................ 5-1

5.1.1 Fairing Separation Sequence ................................................................. 5-1

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5.1.2 Payload Dynamic Design Envelope ....................................................... 5-15.1.3 Payload Access Door ............................................................................ 5-1

5.2 Payload Mechanical Interface and Separation System ....................................... 5-15.2.1 Standard Non-Separating Mechanical Interface..................................... 5-15.2.2 Standard Separating Mechanical Interface............................................. 5-6

5.3 Payload Electrical Interfaces .............................................................................. 5-65.3.1 Umbilical Interfaces .............................................................................. 5-65.3.2 Payload Auxiliary Power ..................................................................... 5-105.3.3 Payload Command and Control .......................................................... 5-105.3.4 Payload Status Monitoring ................................................................... 5-105.3.5 Payload Pyrotechnic Initiator Driver Unit ............................................ 5-105.3.6 Range Safety Interfaces/Vehicle Flight Termination .............................. 5-115.3.7 Electrical Power .................................................................................. 5-115.3.8 Electrical Dead-Facing ........................................................................ 5-135.3.9 Pre-Separation Electrical Constraints ................................................... 5-135.3.10 Non-Standard Interfaces ..................................................................... 5-13

5.4 Payload Design Constraints ............................................................................. 5-135.4.1 Payload Center of Mass Constraints ..................................................... 5-135.4.2 Final Mass Properties Accuracy ........................................................... 5-145.4.3 Payload EMI/EMC Constraints ............................................................. 5-145.4.4 Payload Stiffness ................................................................................. 5-145.4.5 Payload Propellant Slosh ..................................................................... 5-155.4.6 Customer Separation System Shock Constraints .................................. 5-155.4.7 System Safety Constraints .................................................................... 5-15

5.5 Carrier Aircraft Interfaces ................................................................................ 5-155.5.1 Payload Services ................................................................................. 5-155.5.2 Payload Support at Launch Panel ........................................................ 5-16

6.0 MISSION INTEGRATION ............................................................................................ 6-16.1 Mission Management Structure ......................................................................... 6-1

6.1.1 Orbital Mission Responsibilities .............................................................. 6-26.1.1.1 Pegasus Program Management ............................................. 6-26.1.1.2 Pegasus Mission Management .............................................. 6-26.1.1.3 Pegasus Mission Engineering ................................................ 6-26.1.1.4 Pegasus Mechanical Engineering......................................... 6-26.1.1.5 Pegasus Engineering Support ................................................ 6-26.1.1.6 Pegasus Launch Site Operations ........................................... 6-36.1.1.7 Pegasus Systems Safety ........................................................ 6-3

6.2 Mission Integration Process .............................................................................. 6-36.2.1 Mission Teams ...................................................................................... 6-36.2.2 Integration Meetings ............................................................................. 6-36.2.3 Readiness Reviews ................................................................................ 6-4

6.3 Mission Planning and Development ................................................................. 6-46.3.1 Baseline Mission Cycle ......................................................................... 6-4

6.4 Interface Design and Configuration Control ...................................................... 6-6

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iii

6.5 Safety ............................................................................................ 6-66.5.1 System Safety Requirements .................................................................. 6-66.5.2 System Safety Documentation ............................................................... 6-76.5.3 Safety Approval Process ........................................................................ 6-7

7.0 GROUND AND LAUNCH OPERATIONS.................................................................... 7-17.1 Pegasus/Payload Integration Overview .............................................................. 7-17.2 Ground and Launch Operations ....................................................................... 7-1

7.2.1 Launch Vehicle Integration .................................................................... 7-17.2.1.1 Integration Sites ................................................................... 7-17.2.1.2 Vehicle Integration and Test Activities .................................. 7-3

7.2.2 Payload Processing ............................................................................... 7-47.2.2.1 Ground Support Services ..................................................... 7-47.2.2.2 Payload to Pegasus Integration ............................................. 7-47.2.2.2.1 Pre-Mate Interface Testing .................................................... 7-47.2.2.2.2 Payload Mating and Verification ........................................... 7-57.2.2.2.3 Final Processing and Fairing Close-Out ................................ 7-57.2.2.2.4 Payload Propellant Loading .................................................. 7-5

7.2.3 Launch Operations ................................................................................. 7-57.2.3.1 Orbital Carrier Aircraft Mating ............................................. 7-57.2.3.2 Pre-Flight Activities .............................................................. 7-67.2.3.3 Launch Control Organization............................................... 7-67.2.3.4 Flight Activities .................................................................... 7-77.2.3.5 Abort/Recycle/Return-to-Base Operations ............................ 7-7

8.0 DOCUMENTATION ............................................................................................ 8-18.1 Interface Products and Schedules ..................................................................... 8-18.2 Mission Planning Documentation ..................................................................... 8-18.3 Mission-Unique Analyses ................................................................................. 8-1

8.3.1 Trajectory Analysis ................................................................................ 8-18.3.2 Guidance, Navigation and Control Analyses ......................................... 8-28.3.3 Coupled Loads Analysis ........................................................................ 8-28.3.4 Payload Separation Analysis .................................................................. 8-28.3.5 RF Link and Compatibility Analyses ...................................................... 8-28.3.6 Mass Properties Analysis and Mass Data Maintenance .......................... 8-28.3.7 Power System Analysis .......................................................................... 8-28.3.8 Fairing Analyses .................................................................................... 8-38.3.9 Mission-Unique Software ...................................................................... 8-38.3.10 Post-Launch Analysis ............................................................................ 8-3

8.4 Interface Design and Configuration Control ...................................................... 8-3

9.0 SHARED LAUNCH ACCOMMODATIONS .................................................................. 9-19.1 Load-Bearing Spacecraft ................................................................................... 9-19.2 Non Load-Bearing Spacecraft ........................................................................... 9-2

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10.0 SHARED LAUNCH ACCOMMODATIONS ................................................................ 10-110.1 Additional Fairing Access Doors ..................................................................... 10-110.2 Alternative Integration Sites ............................................................................ 10-110.3 Alternative Range Services .............................................................................. 10-110.4 Class 10,000 Fairing Environment ................................................................... 10-110.5 Class 10,000 Payload/Vehicle Integration Environment ................................... 10-110.6 Fairing Internal Surface Cleaning .................................................................... 10-110.7 40-Pin Pass-Through Harness .......................................................................... 10-210.8 Hydrazine Auxiliary Propulsion System .......................................................... 10-210.9 Hydrocarbon Monitoring ................................................................................ 10-210.10 Instrument Purge System................................................................................. 10-210.11 Load Isolation System ..................................................................................... 10-210.12 Low Tip-Off Rate Payload Attach Fittings ......................................................... 10-210.13 Downrange Telemetry Support ........................................................................ 10-210.14 Payload Connector Covers .............................................................................. 10-410.15 Payload Fit Check Support .............................................................................. 10-410.16 Payload Propellant Loading ............................................................................ 10-410.17 Pegasus Separation System Test Unit ............................................................... 10-410.18 Round-the-Clock Payload Support .................................................................. 10-410.19 Serial Telemetry Interface ................................................................................ 10-410.20 Spin Stabilization Above 16.7 RPM ................................................................ 10-410.21 Stage 2 Onboard Camera ............................................................................... 10-510.22 State Vector Transmission From Pegasus .......................................................... 10-510.23 Thermal Coated Forward Separation Ring ....................................................... 10-5

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

APPENDIX B Electrical Interface Connectors ........................................................................B-11.0 Wiring ............................................................................................B-12.0 Connectors ............................................................................................B-13.0 Non-Standard Interfaces ............................................................................................B-1

APPENDIX C VAFB Vehicle Assembly Building Capabilities ................................................ C-11.0 Ground Support Services ........................................................................................... C-12.0 Payload Servicing Areas ........................................................................................... C-13.0 Available Ground Support Equipment ......................................................................... C-14.0 Payload Work Areas ........................................................................................... C-2

APPENDIX D Launch Range Information .............................................................................. D-11.0 Introduction ........................................................................................... D-12.0 Range Safety ........................................................................................... D-1

2.1 Trajectory Analysis .......................................................................................... D-12.2 Area Clearance and Control ............................................................................ D-1

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2.3 Range Safety Displays ..................................................................................... D-12.4 Flight Termination System ............................................................................... D-12.5 FTS Controllers ........................................................................................... D-1

3.0 Telemetry ........................................................................................... D-14.0 Communications ........................................................................................... D-1

4.1 Air to Ground ........................................................................................... D-14.2 Voice Nets ........................................................................................... D-1

5.0 Control Center ........................................................................................... D-16.0 Data Requirements ........................................................................................... D-2

6.1 Realtime Data ........................................................................................... D-26.2 Video Distribution System ............................................................................... D-26.3 Recording ........................................................................................... D-26.4 IRIG Timing ........................................................................................... D-26.5 Weather Forecasts ........................................................................................... D-2

7.0 Optional Launch Ranges ........................................................................................... D-2

APPENDIX E Pegasus Flight History ...................................................................................... E-1

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vi

Figure 2-1 Pegasus XL Just After Release............................................................................ 2-1Figure 2-2 Expanded View of Pegasus XL Configuration.................................................... 2-2Figure 2-3 Principal Dimensions of Pegasus XL (Reference Only) ..................................... 2-3Figure 2-4 Typical Pegasus XL Motor Characteristics in Metri (English) Units .................... 2-4Figure 2-5 Typical Attitude and Guidance Modes Sequence ............................................. 2-5Figure 3-1 Pegasus XL Mission Profile to 741 km (400 nmi) Circular, Polar

Orbit With a 227 kg (501 lbm) Payload ............................................................ 3-1Figure 3-2 Pegasus XL With HAPS Mission Profile to a 741 km (400 nmi) Circular,

Polar Orbit With a 251 kg (554 lbm) Payload................................................... 3-2Figure 3-3 Pegasus XL Performance Capability .................................................................. 3-3Figure 3-4 Typical and Recent Pegasus Orbital Accuracy.................................................. 3-4Figure 3-5 Typical and Recent Orbital Accuracy ............................................................... 3-4Figure 4-1 Factors of Safety for Payload Design and Test ................................................... 4-1Figure 4-2 Payload Testing Requirements .......................................................................... 4-2Figure 4-3 Pegasus Design Limit Load Factors ................................................................... 4-2Figure 4-4 Pegasus XL 3-Sigma High Maximum Acceleratin as a Function

of Payload Weight ............................................................................................ 4-3Figure 4-5 Pegasus Net C.G. LoadFactor Predictions ......................................................... 4-3Figure 4-6 Drop Transient Design Limit Load Environment ............................................... 4-4Figure 4-7 Payload Interface Random Vibration Specification ........................................... 4-4Figure 4-8 Shock at the Base of the Payload ...................................................................... 4-5Figure 4-9 Payload Acoustic Environment ......................................................................... 4-6Figure 4-10 Payload Thermal and Humidity Environment ................................................... 4-7Figure 4-11 Pegasus XL Predicted Worst-Case Payload Fairing Inner Surface

Temperatures During Ascent to Orbit ............................................................... 4-7Figure 4-12 Pegasus XL RF Emitters and Receivers .............................................................. 4-8Figure 4-13 Carrier Aircraft RF Emitters and Receivers ........................................................ 4-8Figure 4-14 Western Range Worst Case Composite Electromagnetic Environment .............. 4-9Figure 4-15 Worse Case Composite Electromagnetic Environment ...................................... 4-9Figure 4-16 Typical Pre-Separation Payload Pointing and Spin Rate Accuracy .................. 4-10Figure 5-1 Payload Fairing Dynamic Envelope with 97 cm (38 in) Diameter

Payload Interface ............................................................................................ 5-2Figure 5-2 Payload Fairing Dynamic Envelope with Optional Hydrazine Auxiliary

Propulsion System (HAPS) and 97 cm (38 in) DiameterPayload Interface ............................................................................................ 5-3

Figure 5-3 Payload Fairing Access Door Placement Zone ................................................. 5-4Figure 5-4 Non-Separable Payload Mechanical Interface .................................................. 5-5Figure 5-5 97 cm (38 in) Separable Payload Interface ....................................................... 5-7Figure 5-6 59 cm (23 in) Separable Payload Interface ....................................................... 5-8

LIST OF FIGURES

FIGURE PAGE



Figure 5-7 43 cm (17 in) Separable Payload Interface ....................................................... 5-9Figure 5-8 Payload Separation Velocities Using the Standard Separation System ............ 5-10Figure 5-9 Pegasus Payload Electrical Interface ............................................................... 5-11Figure 5-10 Pegasus/Spacecraft Electrical Connectors and Associated

Electrical Harnesses ........................................................................................ 5-12Figure 5-11 Pegasus/Spacecraft Pyrotechnic Connectors and Associated

Electrical Harnesses ........................................................................................ 5-12Figure 5-12 Payload Mass vs. Axial c.g. Locatin on X Axis ................................................ 5-13Figure 5-13 Payload Mass Property Measurement Error Tolerances ................................... 5-13Figure 5-14 Detailed RCS Dead Band Zone ...................................................................... 5-14Figure 5-15 Pegasus/OCA Interface Details ....................................................................... 5-15Figure 6-1 Mission Integration Management Structure ....................................................... 6-1Figure 6-2 Summary of Typical Working Groups .............................................................. 6-4Figure 6-3 Typical Mission Cycle ...................................................................................... 6-5Figure6-4 Applicable Safety Requirements ....................................................................... 6-7Figure 6-5 Safety Approval Process ................................................................................... 6-8Figure 7-1 Typical Processing Flow ................................................................................... 7-1Figure 7-2 Typical Pegasus Integration and Test Schedule ................................................. 7-1Figure 7-3 Orbital Carrier Aircraft Hot Pad Area at VAFB .................................................. 7-3Figure 7-4 Pegasus Integration ........................................................................................... 7-3Figure 7-5 Typical Pegasus Launch Checklist Flow ........................................................... 7-6Figure 8-1 Documentation Produced by Orbital for Commercial Pegasus

Launch Services ............................................................................................ 8-1Figure 8-2 Documentation Required by Orbital for Commercial Pegasus

Launch Services ............................................................................................ 8-1Figure 9-1 Load-Bearing Spacecraft Configuration ............................................................ 9-1Figure 9-1 Dual Payload Attach Fitting Configuration ....................................................... 9-2Figure 10-1 Hydrazine Auxiliary Propulsion System (HAPS).............................................. 10-3Figure B-1 Standard Payload Electrical Connections ..........................................................B-1Figure B-2 Payload Interface Connector Pin Assignments for P-65/J-2 Connector .............. B-1Figure C-1 The Vandengerg Vehicle Assembly Building General Layout .......................... C-2Figure D-1 Optional Launch Ranges and Achievable Inclinations ..................................... D-2Figure E-1 Pegasus Rollout ............................................................................................ E-1Figure E-2 Pegasus Launch Locations ................................................................................ E-2

LIST OF FIGURES(CONTINUED)

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Pegasus User's Guide Glossary

A Amperes

AACS Airborne Air Conditioning System

ac Alternating Current

A/C Air Conditioning

AFB Air Force Base

AIT Assembly and Integration Trailer

amps Amperes

ARAR Accident Risk Asessment Report

ARO After Receipt of Order

ASE Airborne Support Equipment

ATP Authority to Proceed

AWG American Wire Gauge

C Centigrade

C/CAM Collision/Contamination Avoidance

Maneuver

CCB Configuration Control Board

CDR Critical Design Review

CFR Code of Federal Regulations

c.g. Center of Gravity

c.m. Center of Mass

cm Centimeter

dB Decibels

dc Direct Current

deg Degrees

DFRF Dryden Flight Research Facility

DoD Department of Defense

DoT Department of Transportation

DPDT Double Pole, Double Throw

EGSE Electrical Ground Support Equipment

EICD Electrical Interface Control Document

EMC Electromagnetic Compatibility

EME Electromagnetic Environment

EMI Electromagnetic Interference

ER Eastern Range (USAF)

F Fahrenheit

FAA Federal Aviation Administration

FAR Federal Acquisition Regulation

fps Feet Per Second

FRR Flight Readiness Review

ft Feet

FTS Flight Termination System

g Gravity

GCL Guidance and Control Lab

GN2 Gaseous Nitrogen

GN&C Guidance, Navigation, and Control

GPS Global Positioning System (NAVSTAR)

Grms Gravity Root Mean Squared

GSE Ground Support Equipment

h Height

HAPS Hydrazine Auxiliary Propulsion System

HEPA High Efficiency Particulate Air

HF High Frequency

HVAC Heating, Ventilating, and Air

Conditioning

H/W Hardware

Hz Hertz

ICD Interface Control Document

IEEE Institute of Electrical and Electronic

Engineers

ILC Initial Launch Capability

IMU Inertial Measurement Unit

in Inch

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INS Inertial Navigation System

ISO International Standardization

Organization

kbps Kilobits per Second

kg Kilograms

km Kilometers

KMR Kwajalein Missile Range

kPa Kilo Pascal

L- Time Prior to Launch

L+ Time After Launch

lbf Pound(s) of Force

lbm Pound(s) of Mass

LOWG Launch Operations Working Group

LPO Launch Panel Operator

LRR Launch Readiness Review

LSC Linear Shaped Charge

m Meters

M Mach

mA Milliamps

MDL Mission Data Load

MHz MegaHertz

MICD Mechanical Interface Control

Document

MIL-STD Military Standard

MIWG Mission Integration Working Group

mm Millimeter

MRR Mission Readiness Review

ms Millisecond

MSD Mission Specification Document

MSPSP Missile System Prelaunch Safety

Package

MUX Multiplexer

m/s Meters Per Second

N2 Nitrogen

N Newtons

N/A Not Applicable

NRTSim Non Real Time Simulation

nm Nautical Miles

NTE Not To Exceed

OASPL Overall Sound Pressure Level

OCA Orbital Carrier Aircraft

OD Operations Directive

OR Operations Requirements Document

Orbital Orbital Sciences Corporation

PDR Preliminary Design Review

PDU Pyrotechnic Driver Unit

P/L Payload

PLF Payload Fairing

POST Program to Optimize Simulated

Trajectories

PPWR P Power

PRD Program Requirements Document

psf Pounds Per Square Foot

psi Pounds Per Square Inch

PSP Program Support Plan

PSSTU Pegasus Separation System Test Unit

PTRN P Turn

PTS Power Transfer Switch

PWP Pegasus Work Package

QA Quality Assurance

RCS Reaction Control System

RF Radio Frequency

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rpm Revolutions Per Minute

RTB Return to Base

RSS Root Summed Squared

S&A Safe & Arm

scfm Standard Cubic Feet Per Minute

sec Second(s)

SIXDOF Six Degree-of-Freedom

S/N Serial Number

S/W Software

SWC Soft Walled Cleanroom

TLM Telemetry

T.O. Take-Off

TT&C Telemetry, Tracking & Commanding

TVC Thrust Vector Control

UDS Universal Documentation System

UFS Ultimate Factory of Safety

USAF United States Air Force

V Volts

VAB Vehicle Assembly Building

VAFB Vandenberg Air Force Base

VDC Volts Direct Current

VHF Very High Frequency

VSWR Voltage Standing Wave Ratio

WFF Wallops Flight Facility

WR Western Range (USAF)

XL Extended Length (Pegasus)

YFS Yield Factor of Safety

Release 5.0 August 2000 xi

Section 1.0—Introduction

Release 5.0 August 2000 1-1

Pegasus User's Guide Section 1.0—Introduction

On August 10, 1989 Orbital Sciences Corporation(Orbital) rolled out the first commerciallydeveloped space launch vehicle for providingsatellites to low earth orbit (see Figure 1-1). Overthe past ten years, the “winged rocket” known asPegasus has proven to be the most successful inits class, placing 70 satellites in orbit with 29launches.

This Pegasus User's Guide is intended tofamiliarize mission planners with the capabilitiesand services provided with a Pegasus launch.

The Pegasus XL was developed as an increasedperformance design evolution from the originalPegasus vehicle to support NASA and the USAFperformance requirements and is now thebaseline configuration for all commercial Pegasuslaunches.

Pegasus is a mature and flight proven smalllaunch system that has achieved consistentaccuracy and dependable performance. ThePegasus launch system has achieved a highdegree of reliability through its significant flightexperience.

Pegasus offers a variety of capabilities that areuniquely suited to small spacecraft. Thesecapabilities and features provide the smallspacecraft customer with greater mission utilityin the form of:

• A range of custom payload interfaces andservices to accommodate unique smallspacecraft missions;

• Payload support services at the PegasusVehicle Assembly Building at VandenbergAFB;

• Horizontal payload integration;

• Shared payload launch accommodations formore cost effective access to space as DualLaunches;

• Portable air-launch capability from worldwidelocations to satisfy unique missionrequirements; and

• Fast, cost-effective and reliable access tospace.

The mobile nature of Pegasus allows Orbital tointegrate the spacecraft to the Pegasus XL in ourintegration facility, the Vehicle Assembly Building(VAB), located at Vandenberg Air Force Base(VAFB), CA and ferry the launch-ready system toa variety of launch ranges. Pegasus has launchedfrom a number of launch locations worldwide(see Figure 1-2).

The unique mobile capability of the Pegasuslaunch system provides flexibility and versatilityto the payload customer. The Pegasus launchvehicle can accommodate integration of thespacecraft at a customer desired location as wellas optimize desired orbit requirements based onthe initial launch location. In 1997, after finalbuild up of the rocket at the VAB, Pegasus wasmated to the Orbital Carrier Aircraft (OCA) andferried to Madrid, Spain to integrate Spain’sMINISAT-01 satellite. Following integration ofthe satellite, Pegasus was then ferried to theisland of Gran Canaria for launch. The successfullaunch of Spain’s MINISAT-01 satellite provedout Pegasus’ ability to accommodate the payloadprovider’s processing and launch requirementsat locations better suited to the customer ratherthan the launch vehicle. This unprecedentedlaunch vehicle approach is an example ofPegasus’s way of providing customer orientedlaunch service.

In the interest of continued process improvementand customer satisfaction, the Pegasus Programsuccessfully completed a one year effort of ISOFigure 1-1. Pegasus Rollout.

PEG001


Pegasus User's Guide Section 1.0—Introduction

9001 certification. In July 1998, Orbital’s LaunchSystems Group was awarded this internationallyrecognized industry benchmark for operating aquality management system producing a qualityproduct and service. In the true spirit of ISO9001, this level of quality is not only achieved,but must be maintained. To this end, the LaunchSystems Group has successfully passed eachsemi-annual audit since the award in 1998.

Pegasus is a customer oriented and responsivelaunch vehicle system. From Pegasus’ commercialheritage comes the desire to continually addressthe payload customer market to bestaccommodate its needs. The Pegasus launchvehicle system has continually matured andevolved over its ten year history. This ability anddesire to react to the customer has produced thesingle most successful launch vehicle in its class.To ensure our goal of complete customersatisfaction, a team of managers and engineers isassigned to each mission from “contract award topost-flight report”. This dedicated team is

Western Range

70° to 130° Inclination

Kwajalein Atoll0° to 10° Inclination

Equator

Eastern Range28° to 50° Inclination

Canary IslandsLaunch PointMobile Range25° Inclination(Retrograde)

Wallops Flight Facility30° to 65° Inclination

Figure 1-2. Pegasus Launch Locations.PEG002

Alcantara Launch Center0° to 90° Inclination

TorrejonAir Base

committed to providing the payload customer100% satisfaction of mission requirements.

Each Pegasus mission is assigned a mission teamled by a Mission Manager and a Mission Engineer.The mission team is responsible for missionplanning and scheduling, launch vehicleproduction coordination, payload integrationservices, systems engineering, mission-peculiardesign and analysis, payload interface definition,range coordination, launch site processing andoperations. The mission team is responsible forensuring all mission requirements have beensatisfied.

Section 2.0—Pegasus XL Vehicle Description


Pegasus User's Guide Section 2.0—Pegasus XL Vehicle Description

2.1 Pegasus XL Vehicle Description

As discussed in Section 1.0, Pegasus continues toevolve in response to customer requirements.The initial configuration of Pegasus (referred toas the Standard Pegasus) was modified to provideincreased performance and vehicleenhancements. The last of the Pegasus Standardlaunch vehicles is expected to be launched bythe end of 2000, therefore, this Pegasus User’sGuide is dedicated to the discussion of thePegasus XL configuration, capabilities, andassociated services.

Pegasus XL is a winged, three-stage, solid rocketbooster which weighs approximately 23,130 kg(51,000 lbm) and measures 16.9 m (55.4 ft) inlength and 1.27 m (50 in) in diameter and has awing span of 6.7 m (22 ft). Figure 2-1 shows thePegasus on the Assembly Integration Trailer (AIT).Pegasus is lifted by the Orbital Carrier Aircraft(OCA) to a level flight condition of about 11,900m (39,000 ft) and Mach 0.80. Five seconds afterrelease from the OCA stage 1 motor ignitionoccurs. The vehicle's autonomous guidanceand flight control system provide the guidancenecessary to insert payloads into a wide range oforbits.

Figure 2-2 shows an expanded view of the PegasusXL configuration. The Pegasus Vehicle designcombines state-of-the-art, flight-proventechnologies, and conservative design marginsto achieve performance and reliability at reduced

cost. The vehicle incorporates eight majorelements:

• Three solid rocket motors;

• A payload fairing;

• An avionics assembly;

• A lifting wing;

• Aft skirt assembly including three movablecontrol fins; and

• A payload interface system.

Pegasus also has an option for a liquid propellantfourth stage, HAPS (see Section 10). Figure 2-3illustrates Pegasus XL's principle dimensions.

2.1.1 Solid Rocket Motors

The three solid rocket motors were designed andoptimized specifically for Pegasus and includefeatures that emphasize reliability,manufacturability, and affordability. The designwas developed using previously flight-provenand qualified materials and components.Common design features, materials, andproduction techniques are applied to all threemotors to maximize cost efficiency and reliability.These motors are fully flight-qualified. Typicalmotor characteristics are shown in Figure 2-4.

2.1.2 Payload Fairing

The Pegasus payload fairing consists of twocomposite shell halves, a nose cap integral to ashell half, and a separation system. Each shellhalf is composed of a cylinder and ogive sections.The two halves are held together with two titaniumstraps along the cylinder and a retention bolt inthe nose. A cork and Room TemperatureVulcanizing (RTV) Thermal Protection System(TPS) provides protection to the graphitecomposite fairing structure. The amount of TPSapplied has been determined to optimize fairingperformance and payload environmentalprotection.

The two straps are tensioned using bolts, whichare severed during fairing separation withpyrotechnic bolt cutters, while the retention bolt

Figure 2-1. Pegasus XL on the Assembly and Inte-gration Trailer (AIT).

PEG003



in the nose is released with a pyrotechnicseparation nut. The base of the fairing is separatedwith Orbital's low-contamination frangibleseparation joint. These ordnance events aresequenced for proper separation dynamics. Ahot gas generator internal to the fairing is alsoactivated at separation to pressurize two piston-driven pushoff thrusters. These units, inconjunction with cams, force the two fairinghalves apart. The halves rotate about fall-awayhinges, which guide them away from the satelliteand launch vehicle.

The fairing and separation system were fullyqualified through a series of structural, functional,and contamination ground vacuum tests andhave been successfully flown on all Pegasus XLmissions. Section 5 presents a more detaileddescription of the fairing separation sequenceand the satellite dynamic envelope.

2.1.3 Avionics

The Pegasus avionics system is a digital distributed

processor design that implements recentdevelopments in hardware, software,communications, and systems design. Missionreliability is achieved by the use of simple designs,high-reliability components, high design marginsand extensive testing at the component, subsystemand system level.

The heart of the Pegasus avionics system is amultiprocessor, 32-bit flight computer. The flightcomputer communicates with the InertialMeasurement Unit (IMU), the launch panelelectronics on the carrier aircraft and all vehiclesubsystems using standard RS-422 digital serialdata links. Most avionics on the vehicle featureintegral microprocessors to perform localprocessing and to handle communications withthe flight computer. This RS-422 architecture iscentral to Pegasus's rapid integration and test, asit allows unit and system-level testing to beaccomplished using commercially availableground support equipment with off-the-shelfhardware.

Interstage

Fin

Wing

Aft SkirtAssembly

Stage 1 Motor

AvionicsSection

PayloadSeparationSystem

Stage 2 Motor

*Stage 3 Motor

PayloadFairing

*Optional 4th Stage Available for Precision Injection

Figure 2-2. Expanded View of Pegasus XL Configuration.PEG004



2.1.4 Flight Termination System

The Pegasus Flight Termination System (FTS)supports ground-initiated command destruct aswell as the capability to sense inadvertent stageseparation and automatically destruct the rocket.The FTS is redundant, with two independent safeand arm devices, receivers, logic units, andbatteries.

2.1.5 Attitude Control Systems

After release from the OCA, the Pegasus attitudecontrol system is fully autonomous. Acombination of open-loop steering and closed-loop guidance is employed during the flight.Stage 1 guidance utilizes a pitch profile optimizedby ground simulations. Stages 2 and 3 guidanceuses an adaptation of an algorithm that was first

Figure 2-3. Principal Dimensions of Pegasus XL (Reference Only).

+X

+Z

Stage 1/Stage 2Second Separation

Side View Stage 1/Stage 2First Separation

+ Pitch

- Pitch

1,081.0425.6

STA. +

127.050.0

Φ

47.018.5

STA. +

190.575.0

STA. +

1,000.5393.9

STA. +

570.0224.4

+X

+Y

Fairing Separation and Fifth Hook

Stage 2/Stage 3Separation

Payload Interface Plane (22" Long Avionics Structure)

Top View Looking Down

- Yaw

+Yaw

STA. +

1,475.5580.9

STA. +

1,741685.4

STA. +

1,354.1533.1

STA. +1,299.7511.7

STA. +

180°

270° 90°

0°

23°

+Y

+Z

Aft View Looking Forward

Dimensions cmin

Note: STA. Reference is a Point in Space 47.0 cm (18.5") Aft of the Stage 1 Nozzle Total Vehicle Length: 1,693.9 cm (666.9")

- Roll

+ Roll

176.569.5

STA. +

281.2110.7

402.9158.6

670.6264.0

PEG006



developed for the Space Shuttle ascent guidance.Attitude control is closed-loop.

The vehicle attitude is controlled by the FinActuator System (FAS) during Stage 1 flight. Thisconsists of electrically actuated fins located atthe aft end of Stage 1. For Stage 2 and Stage 3flight, a combination of electrically activatedThrust Vector Controllers (TVCs) on the Stage 2and Stage 3 solid motor nozzles and a GN2

Reaction Control System (RCS) system locatedon the avionics section, control the vehicleattitude.

Figure 2-5 summarizes the attitude and guidancemodes during a typical flight, although the exactsequence is controlled by the Mission Data Load(MDL) software and depends on mission specificrequirements.

2.1.6 Telemetry Subsystem

The Pegasus XL telemetry system provides realtime health and status data of the vehicle avionicssystem, as well as key information regarding theposition, performance and environment of thePegasus XL vehicle. This data may be used byOrbital and the range safety personnel to evaluatesystem performance.

Pegasus contains two separate telemetry systems.The first provides digital data through telemetrymultiplexers (MUXs) which gather data fromeach sensor, digitize it, then relay the informationto the flight computer. This Pegasus telemetrystream provides data during ground processing,checkout, captive carry, and during launch.During captive carry, Pegasus telemetry isdownlinked to the ground and recorded onboardthe OCA. Some payload telemetry data can beinterleaved with Pegasus data as a non-standardservice. The second system provides analogenvironments data which are transmitted via awideband data link and recorded for post-flightevaluation.

2.1.7 Major Structural Subsystems

2.1.7.1 Wing

The Pegasus wing uses a truncated delta platformwith a double wedge profile. Wing panels aremade of a graphite-faced Nomex-foam sandwich.Channel section graphite spars carry the primarybending loads and half-ribs and reinforcing lay-ups further stabilize the panels and reduce stressconcentrations. The wing central box structurehas fittings at each corner which provide thestructural interface between the Pegasus and theOCA.

Stage 3 MotorOrion 38

Stage 2 MotorOrion 50 XL

Stage 1 MotorOrion 50S XLUnitsParameter

cm (in)

cm (in)

kg (lb)

kg (lb)

kN-sec (lbf-sec)

kPa (psia)

sec

kN (lbf)

Nsec/kg (lbf-sec/lbm)

deg

1,027 (404)

128 (50)

1,369 (3,019)

15,014 (33,105)

43,586 (9,799,080)

7,515 (1,090)

68.6

726 (163,247)

2,846 (295)

NA

311 (122)

128 (50)

416 (918)

3,925 (8,655)

11,218 (2,522,070)

7,026 (1,019)

69.4

196 (44,171)

2,838 (289)

±3

134 (53)

97 (38)

126 (278)

770 (1,697)

2,185 (491,200)

4,523 (656)

68.5

36 (8,062)

2,817 (287)

±3

Figure 2-4. Typical Pegasus XL Motor Characteristics in Metric (English) Units.

Overall Length

Diameter

Inert Weight (1)

Propellant Weight (2)

Total Vacuum Impulse (3)

Average Pressure

Burn Time (3) (4)

Maximum Vacuum Thrust (3)

Vacuum Specific Impulse Effective (5)

TVC Deflection

(1) Including Wing Saddle, Truss, and Associated Fasteners(2) Includes Igniter Propellants

(3) At 21°C (70° F)(4) To 207 kPa (30 psi)(5) Delivered (Includes Expended Inerts)

Notes:

PEG007



2.1.7.2 Aft Skirt Assembly

The aft skirt assembly is composed of the aft skirt,three fins, and the fin actuator subsystem. The aftskirt is an all-aluminum structure of conventionalring and stressed-skin design with machinedbridge fittings for installation of theelectromechanical fin actuators. The skirt issegmented to allow installation around the firststage nozzle. Fin construction is one-piece solidfoam core and wet-laid graphite compositeconstruction around a central titanium shaft.

2.1.7.3. Payload Interface Systems

Multiple mechanical and electrical interfacesystems currently exist to accommodate a varietyof spacecraft designs. Section 5.0 describesthese interface systems. To ensure optimizationof spacecraft requirements, payload specificmechanical and electrical interface systems canbe provided to the payload customer. Payloadmechanical fit checks and electrical interfacetesting with these spacecraft unique interface

systems are encouraged to ensure all spacecraftrequirements are satisfied.

2.2 Orbital Carrier Aircraft

Orbital furnishes and operates the Orbital CarrierAircraft (OCA). After integration at Orbital’sWest Coast integration site at VAFB, the OCA canprovide polar and high-inclination launchesutilizing the tracking, telemetry, and command(TT&C) facilities of the WR. The OCA canprovide lower inclination missions from the EastCoast using either the NASA or ER TT&C facilities,as well as equatorial missions from the KwajaleinAtoll or Alcantara, Brazil. The OCA is madeavailable for mission support on a priority basisduring the contract-specified launch window.

The unique OCA-Pegasus launch systemaccommodates two distinctly different launchprocessing and operations approaches for non-VAFB launches. One approach (used by themajority of payload customers) is to integrate thePegasus and payload at the VAB and then ferry

Guidance ModeMajor Phase

Free Drop

Stage 1 Flight

Stage 1 Flight

Stage 1 Flight

Stage 1 Flight

Stage 1/2 Coast

Stage 2 Flight

Stage 2/3 Coast

Stage 2/3 Coast

Stage 3

After Stage 3 Burnout

Ignition and Pull Up

Maximum Pitch Up

Pitch Over

Fins Zeroed

S2 Ignition

Maneuver to S3 Ignition Attitude

S3 Ignition

Payload Events as Required

Inertial Euler Angles


Vertical Acceleration Limit


Attitude Hold

Attitude Hold

Commanded Attitude

Attitude Hold

Commanded Attitude

Commanded Attitude

Commanded as Required

Figure 2-5. Typical Attitude and Guidance Modes Sequence.

Minor Phase Attitude Mode

None

Nominal Trajectory

Nominal Trajectory

Nominal Trajectory

None

Begin S2 Powered Explicit GuidanceSolution

Closed Loop PoweredExplicit Guidance

Begin S3 Powered Explicit GuidanceSolution

None

Closed Loop PoweredExplicit Guidance

None

PEG008

Mission Specific Events Tailored to Payload Requirements

Fixed Events for All Missions



the integrated Pegasus and payload to anotherlocation for launch. This approach is referred toas a “ferry mission.” The second approach isreferred to as a “campaign mission.” A campaignmission starts with the build up of the Pegasus atthe VAB. The Pegasus is then mated to the OCAat VAFB and then ferried to the integration sitewhere the Pegasus and payload are fully integratedand tested. At this point, the launch may eitheroccur at the integration site or the integratedPegasus and payload may be ferried to anotherlocation for launch.

The OCA also has the capability to ferry Pegasustrans-continentally or trans-oceanically(depending on landing site) to support ferry andcampaign missions.

Section 3.0—General Performance Capability


Pegasus User's Guide Section 3.0—General Performance Capability

This section describes the orbital performancecapabilities of the Pegasus XL vehicle with andwithout the optional Hydrazine AuxiliaryPropulsion System (HAPS) described in Section10. Together these configurations can deliverpayloads to a wide variety of circular and ellipticalorbits and trajectories, and attain a completerange of prograde and retrograde inclinationsthrough a suitable choice of launch points andazimuths. In general, HAPS will provideadditional performance at higher altitudes.

From the Western Range (WR), Pegasus canachieve inclinations between 70° and 130°. Abroader range of inclinations may be achievable,subject to additional analyses and coordinationwith Range authorities. Additionally, lowerinclinations can be achieved through dog-legtrajectories, with a commensurate reduction inperformance. Some specific inclinations withinthis range may be limited by stage impact pointor other restrictions. Other inclinations can besupported through use of Wallops Flight Facility

(WFF), Eastern Range (ER) or other remote TT&Csites. Pegasus requirements for remote sites arelisted in Appendix D.

3.1 Mission Profiles

This section describes circular low earth orbitmission profiles. Performance quotes for non-circular orbits will be provided on a mission-specific basis.

Profiles of typical missions performed by PegasusXL with and without HAPS are illustrated inFigure 3-1 and Figure 3-2. The depicted profilebegins after the OCA has reached the launchpoint, and continues through orbit insertion. Thetime, altitude, and velocity for the major ignition,separation, and burnout events are shown for atypical trajectory that achieves a 741 km (400nm) circular, polar (90° inclination) orbit afterlaunch from WR. These events will vary based onmission requirements.

The typical launch sequence begins with release

Figure 3-1. Pegasus XL Mission Profile to 741 km (400 nmi) Circular, Polar Orbit with a 227 kg (501 lbm)Payload.

PEG005

Third Stage Ignition t = 592 sec h = 398.9 nmi v = 14,864 fps g= 1.5 deg

OrbitalSciencesCorporation

HERCULES

1

L-1011 Drop Launch t = 0 h = 38,000 ft v = 770 fps

First Stage Ignition t = 5 sec h = 37,640 ft v = 1,450 fps

Max q1,500 psf

First Stage Burnout t = 77 sec h = 207,140 ft v = 8,269 fps

Second Stage Ignition t = 95.3 sec h = 288,600 ft v = 7,969 fps

Payload Fairing Separation t = 110.6 sec h = 356,390 ft v = 8,892 fps

Second Stage Burnout t = 168 sec h = 709,070 ft v = 17,809 fps

Second/ThirdStage Coast

Third Stage Burnoutand Orbital Insertion t = 657 sec h = 400 nmi v = 24,770 fps g = 0.0 deg



of Pegasus from the carrier aircraft at an altitudeof approximately 11,900 m (39,000 ft) and aspeed of Mach 0.80. Approximately 5 secondsafter drop, once Pegasus has cleared the aircraft,Stage 1 ignition occurs. The vehicle quicklyaccelerates to supersonic speed while beginninga pull up maneuver. Maximum dynamic pressureis experienced approximately 25 seconds afterignition. At approximately 20-25 seconds, amaneuver is initiated to depress the trajectoryand the vehicle angle of attack quickly approacheszero.

Stage 2 ignition occurs shortly after Stage 1burnout and the payload fairing is jettisonedduring Stage 2 burn as quickly as fairing dynamicpressure and payload aerodynamic heatinglimitations will allow, approximately 110,000 m(361,000 ft) and 112 seconds after drop. Stage 2burnout is followed by a long coast, duringwhich the payload and Stage 3 achieve orbital

altitude. Stage 3 then provides the additionalvelocity necessary to circularize the orbit. Stage3 burnout typically occurs approximately 10minutes after launch and 2,200 km (1,200 nm)downrange of the launch point. Attitude controlduring Stage 2 and Stage 3 powered flight isprovided by the motor Thrust Vector Control(TVC) system for pitch and yaw and by thenitrogen cold gas Reaction Control System (RCS)for roll. The RCS also provides control about allthree axes during coast phases of the trajectory.

3.2 Performance Capability

Performance capabilities to various orbits for thePegasus XL are illustrated in Figure 3-3. Theseperformance data were generated using theProgram to Optimize Simulated Trajectories(POST), which is described below. Preciseperformance capabilities to specific orbits areprovided per the timeline shown in Section 8.0.

HERCULES

1

L-1011 Drop Launch t = 0 h = 39,000 ft v = 797 fps

First Stage Ignition t = 5 sec h = 38,690 ft v = 800 fps

Max q1,500 psf

First Stage Burnout t = 76 sec h = 141,096 ft v = 8,000 fps

Second Stage Ignition t = 97.48 sec h = 228,000 ft v = 9,170 fps

Payload Fairing Separation t = 136 sec h = 343,000 ft v = 13,520 fps

Second Stage Burnout t = 167.12 sec h = 454,000 ft v = 19,000 fps

Second/ThirdStage Coast

Third Stage Burnout t = 458 sec h = 949,000 ft v = 25,790 fps

Third Stage Ignition t = 391.04 sec h = 935,000 ft v = 182,020 fps

Stage 3/HAPSSeparation t = 517.68 sec h = 954,000 ft v = 25,780 fps

Figure 3-2. Pegasus XL With HAPS Mission Profile to a 741 km (400 nmi) Circular, Polar Orbit With a 251 kg (554 lbm) Payload.

PEG009



3.3 Trajectory Design Optimization

Orbital designs a unique mission trajectory foreach Pegasus flight to maximize payloadperformance while complying with the satelliteand launch vehicle constraints. Using POST, adesired orbit is specified and a set of optimizationparameters and constraints are designated.Appropriate data for mass properties,aerodynamics, and motor ballistics are input.POST then selects values for the optimizationparameters that target the desired orbit withspecified constraints on key parameters such asangle of attack, dynamic loading, payloadthermal, and ground track. After POST has beenused to determine the optimum launch trajectory,a Pegasus-specific six degree of freedomsimulation program is used to verify trajectoryacceptability with realistic attitude dynamics,including separation analysis on all stages.

Figure 3-3. Pegasus XL Performance Capability.

500

450

400

350

300

250

200

150

100

50

200 600400 800 1,000 1,200 1,400 km

(1)Requires VAFB Waiver

Pay

load

Cap

abili

ty (k

g)

Pay

load

Cap

abili

ty (l

bm)

Circular Orbit Altitude

0

400

500

600

700

800

900

1,000

1,100

300

200

100

0

• Drop Conditions: 11,900 m (39,000 ft) Mach 0.80

nmi200 300 400 500 600 700

• Entire Mass of the Separation System Is Bookkept on the Launch Vehicle Side• 67 m/sec (220 ft/sec) Guidance Reserve Maintained• Fairing Jettison at .48 Pa (0.01 lbf/ft2)

28.5° Orbit from Eastern Range38° Orbit from Wallops Flight Facility60°(1) Orbit from Western Range70° Orbit from Western Range90° Orbit from Western Range98° Orbit from Western Range

PEG010

3.4 Injection Accuracy

Figure 3-4 provides estimates of 3-sigma orbitalinjection errors for a 227 kg (501 lbm) payload toa 741 km (400 nm), circular, 90° inclinationreference orbit. These errors are dominated byerrors of the final propulsive stage. In general,the insertion apse experiences smaller errorsthan the non-insertion apse.

Orbital injection errors are inherently missionspecific for solid stage vehicles. In generalhowever, for most missions, insertion accuracieswill not be radically different than the valuesquoted in Figure 3-4. Total orbital altitude errorsare dominated by errors associated with the finalpropulsive stage. Several factors affect orbitalaccuracy directly. Payload masses have the largesteffect because they affect the velocity errorresulting from a given motor impulse error. Lighterpayloads will net greater non-insertion apse errors



than a heavy payload for a given target.Additionally the choice of guidance strategy tomeet particular mission requirements can alsoaffect orbital errors.

3.4.1 Actual Pegasus Injection Accuracies

Figure 3-5 shows actual Pegasus orbital injectionaccuracies for missions in 1996 and 1997 havebeen consistently within one sigma bounds. Asa benchmark, on a typical Pegasus mission, onesigma corresponds to an insertion apse accuracyof ±5 km and a non-insertion apse accuracy of±30 km. Orbital inclination accuracies have alsobeen well within one sigma. Typical inclinationerrors are within ±0.05°.

Accuracies are highly mission-specific,depending on payload mass, targeted orbit, andthe particular guidance strategy adopted for themission. In particular, light payloads and highorbits experience increased injection error.Conversely, heavy payloads and low orbitsexperience reduced injection error. Preliminaryand final mission specific orbital dispersions areprovided in the Preliminary and Final MissionAnalyses.

3.4.2 Error-Minimizing Guidance Strategies

Pegasus motor performance, mass properties andguidance system are understood very well due tolarge amount of actual flight experience to date.This historical record has enabled the PegasusProgram to update the vehicle models toaccurately predict mission performance.

In order to assure that even a 3σ low-performancePegasus will achieve the required orbit, Pegasustrajectories include a 54 m/sec (180 ft/sec)

±19 km±15 km

±90 km±15 km

±0.15°±0.08°

InsertionApse Error

Non-InsertionApse Error

InclinationErrorConfiguration

Figure 3-4. 3-Sigma Injection Accuracies TypicalPegasus XL Missions.

Pegasus XLPegasus XLwith HAPS P

EG

070

Figure 3-5. Typical and Recent Pegasus Orbital Accuracy.

50

40

30

20

10

0

-10

-20

-30

-40

-50 -40 -30 -20 -10 0 10 20 30 40 50

Apo

gee

Del

ta F

rom

Tar

get (

km)

Perigee Delta From Target (km) * HAPS Missions

-50

PEG111

F16 (Apogee = -98 km, Perigee = -10 km)

F14

F17

F24

F22*

F18*

F15

F10

F12

F29

F13

F20

F25

F11

F23*

F19* F21

F26 (Apogee = -58 km, Perigee = -1 km)

LR-81 INSF10-REX IIF11-MSTI-3F12-TOMSF13-FASTF14-SAC-B/HETEF15-MINISATF16-ORBVIEW-2F17-FORTEF18-STEP-4F19-ORBCOMM-1F20-SNOE/BATSATF21-TRACE

LN-100 INSF22-ORBCOMM-2F23-ORBCOMM-3F24-SCD-2F25-SWASF26-WIREF27-TERRIERS/MUBLCOMF28-ORBCOMM-4F29-TSX-5

-0.04-0.04-0.03-0.02-0.02-0.04+0.07-0.03-0.02+0.02+0.01-0.07

0.010.00

-0.01-0.09-0.03-0.03

+0.005-0.05

Inclination Delta From Target (Deg)

F27M*

F28*F27T*



guidance reserve. Pegasus software allows avariety of error-minimizing guidance strategiesto be used with this reserve. These strategies fallinto three basic categories:

(1) Minimize Insertion Errors. Using this strategy,the guidance system scrubs off excess energyvia out of plane turning during Stage 2 and 3burns and modifying the coast durationbetween Stage 2 and 3 burns. This strategyresults in the smallest possible insertion errorsfor both apogee and perigee.

(2) Maximize Apogee Altitude. Using thisstrategy, all excess velocity is conserved inorder to maximize velocity at insertion. Thisallows the customer to take advantage of theguidance reserve by increasing the expectedapogee altitude while maintaining a preciseperigee altitude.

(3) Some Combination of (1) and (2). Options 1and 2 are the two endpoints of a spectrum ofpotential guidance strategies. A third optioncan target a particular insertion velocity higherthan the 3-DOF nominal capability, but lowerthan the vehicle's 3σ high capability. Usingthis "hybrid" approach, if the desired apogeealtitude corresponds to an insertion velocitywhich is "X" m/sec higher than the nominal 3-DOF insertion velocity, then the vehicle willnot scrub energy unless an excess of greaterthan "X" m/sec above the nominal 3-DOFvalue is achieved. This strategy results in anapogee distribution where the mean valuefalls between the results from options 1 and2. The total apogee dispersions will be largerthan those resulting from option 1, but smallerthan those from option 2.

3.5 Collision/Contamination AvoidanceManeuver

Following orbit insertion, the Pegasus Stage 3RCS or HAPS will perform a series of maneuverscalled a Collision/Contamination AvoidanceManeuver (C/CAM). The C/CAM minimizes bothpayload contamination and the potential forrecontact between Pegasus hardware and the

separated payload. It also depletes all remainingnitrogen and/or hydrazine.

Orbital will perform a recontact analysis for postseparation events. Orbital and the payloadcontractor are jointly responsible fordetermination of whether a C/CAM is required.

A typical C/CAM consists of the following steps:

1) At payload separation +3 seconds, the launchvehicle performs a 90° yaw maneuverdesigned to direct any remaining State 3motor impulse in a direction which willincrease the separation distance betweenthe two bodies.

2) At payload separation +300 seconds, thelaunch vehicle begins a “crab-walk”maneuver. This maneuver, performedthrough a series of RCS thruster firings, isdesigned to impart a small amount of deltavelocity in the negative velocity vectordirection, increasing the separation velocitybetween the payload and the third stage ofthe Pegasus. The maneuver is terminatedapproximately 600 seconds after separation.

3) Following the completion of the C/CAMmaneuver as described above, the RCS valvesare opened and the remaining gas is expelled.

Section 4.0—Payload Environments


Pegasus User's Guide Section 4.0—Payload Environments

This section describes the payload environmentsexperienced through the ground, captive carryand flight mission phases. In most cases bothdesign limit loads and measured flight data arecharacterized. These limit loads encompass theenvironments imposed by the XL and HAPSconfigured vehicles and by the Orbital CarrierAircraft (OCA).

4.1 Design Loads

The primary support structure for the spacecraftshall possess sufficient strength, rigidity, andother characteristics required to survive the criticalloading conditions that exist within the envelopeof handling and mission requirements, includingworst case predicted ground, flight, and orbitalloads. It shall survive those conditions in amanner that assures safety and that does notreduce the mission success probability. Theprimary support structure of the spacecraft shallbe electrically conductive to establish a singlepoint electrical ground. Spacecraft design loadsare defined as follows:

• Design Limit Load — The maximum predictedground-based, captive carry or powered flightload, including all uncertainties.

• Design Yield Load — The Design Limit Loadmultiplied by the required Yield Factor ofSafety (YFS) indicated in Figure 4-1. Thepayload structure must have sufficient strengthto withstand simultaneously the yield loads,applied temperature, and otheraccompanying environmental phenomenafor each design condition withoutexperiencing detrimental yielding orpermanent deformation.

• Design Ultimate Load — The Design LimitLoad multiplied by the required UltimateFactor of Safety (UFS) indicated in Figure 4-1.The payload structure must have sufficientstrength to withstand simultaneously theultimate loads, applied temperature, and otheraccompanying environmental phenomenawithout experiencing any fracture or otherfailure mode of the structure.

4.2 Payload Testing and Analysis

Sufficient payload testing and/or analysis must beperformed to ensure the safety of ground andaircraft crews and to ensure mission success. Thepayload design must comply with the testing anddesign factors of safety in Figure 4-1 and the FAAregulations for the carrier aircraft listed in CFR14document, FAR Part 25. Ultimate Factors ofSafety shown in Figure 4-1 must be maintainedper Orbital SSD TD-0005. At a minimum, thefollowing tests must be performed:

Structural Integrity — Static loads, sine vibration,or other tests shall be performed that combine toencompass the acceleration load environmentpresented in Section 4.3. Test level requirementsare defined in Figure 4-1.

Random Vibration — Test level requirements aredefined in Figure 4-2.

4.3 Payload Acceleration Environment

Figure 4-3 illustrates the primary accelerationload conditions experienced during a nominalPegasus integration and launch operation usingthe Orbital Carrier Aircraft. The accelerationslisted are design limit loads. The axialaccelerations for each stage at burnout arepresented in Figure 4-4.

Design andTest Options

Figure 4-1. Factors of Safety for Payload Design and Test.

Dedicated Test Article

Proto-Flight Article

Proof Test Each Flight Article

No Static Test

(YFS)

Ultimate

(UFS) UnmannedEvents

(UFS) MannedEvents

Yield

Design Factor of Safety on Limit Loads

TestLevel

1.00

1.25

1.10

1.60

1.25

1.50

1.25

2.00

1.50

1.50

1.50

2.25

UFS

1.25

1.10

N/A

PEG012



4.3.1 Drop Transient Acceleration

The Pegasus has no significant sustainedsinusoidal vibration environments during captivecarry or powered flight. There is a transientacceleration event, which occurs during the dropof the Pegasus from the carrier aircraft. Prior tothe Pegasus separation, the Pegasus/payloadstructure is deformed due to the gravitational pre-load. At drop, the pre-load is suddenly removed.The resulting transient response is dominated bythe Pegasus/Payload first bending mode (8-9Hz). However, higher frequency Pegasus andpayload modes are excited as well. Because ofthe oscillatory nature of the drop transientresponse, which includes rotation of the interfaceplane, significant dynamic amplification of theaccelerations is expected throughout thespacecraft. The mass distribution, stiffness andlength of the primary payload structure greatlyimpact the amplification level. Accurateestimation of the drop transient loading requiresa coupled loads analysis (CLA) which uses Orbitaland customer provided finite element models topredict the drop transient environment. Prior toperforming a CLA, Figure 4-5 can be used toestimate the payload c.g. Net Load Factors (forthe Pegasus Z-axis) and the payload interfaceestimates are shown in Figure 4-6. Load factorsfor other payload interface configurations, or formodified 23” and 38” separation systems (i.e.,

load suppression), require mission specificanalyses for accurate predictions. To minimizecoupling of the payload bending modes with thelaunch vehicle first bending mode, the firstfundamental lateral frequency must be greaterthan 20 Hz, cantilevered from the base of thespacecraft, excluding the spacecraft separationsystem.

4.4 Payload Vibration Environment

Based on flight data taken during OCA captivecarry flights, the in flight random vibration curveshown in Figure 4-7 encompasses the captivecarry vibration environment.

4.4.1 Long Duration Captive Carry

The maximum envelope shown in Figure 4-7 isnot constant during a Pegasus mission. Theactual flight random vibration levels varyconsiderably throughout each phase of thePegasus flight and are typically well below themaximum levels.

4.5 Payload Shock Environment

The maximum shock response spectrum at thebase of the payload from all launch vehicleevents will not exceed the flight limit levels inFigure 4-8.

4.6 Payload Acoustic Environment

The acoustic levels during OCA take-off, captivecarry and powered flight will not exceed theflight limit levels shown in Figure 4-9. The +6dBspectrum is recommended for payload standardacoustic testing to account for fatigue durationeffects.

Flight Limit Level + 6dB

Flight Limit Level

Flight Limit Level + 3dB

Qualification

Acceptance

Protoflight

Random Vibration: theFlight Limit Level IsCharacterized inFigure 4-7

Test Type

Test

Purpose Test Level

Figure 4-2. Payload Testing Requirements.

PEG013

Environment

Figure 4-3. Pegasus Design Limit Load Factors.

Steady-

State

Steady-

State

Steady-

StateN/A

-3.7

See Fig. 4-4

±0.2

±1.0

±1.0

±1.0

±1.0

Quasi-Static*±0.7

±1.0

±1.0

±2.0

Quasi-Static*

Y-Axis (g’s) Z-Axis (g’s)X-Axis (g’s)

N/A

±0.2

±0.2

±0.2

Quasi-Static*+3.6/-1.0

+1.0

±1.0

±2.0

N/A

±2.33

±0.2

±0.2

Taxi, Captive Flight & Abort Landing (Man-Rated)2

Aerodynamic Pull-Up

Stage Burn-Out

Post Stage Burn-OutNotes:

1) Static Equivalent of Mixed Dynamic Environments

2) Dominated by Abort and Ferry Landing Events. PEG017



4.7 Payload Thermal and Humidity Environment

The payload temperature and humidityenvironments are controlled inside the fairingusing the Ground and Airborne Air ConditioningSystems (GACS and AACS). The GACS providesconditioned air to the payload in the VAB, on the

flight line. The AACS is used prior to OCA take-off and during captive carry flight. Theconditioned air enters the fairing at a locationforward of the payload, exits aft of the payloadand is provided up to the time of launch vehicledrop. Baffles are provided at the air conditioninginlet to reduce impingement velocities on the

50

Payload Mass

3-S

igm

a H

igh

Max

imum

Axi

al A

ccel

erat

ion

(G's

)

Figure 4-4. Pegasus XL 3-Sigma High Maximum Acceleration as a Function of Payload Weight.

14.0

13.0

12.0

11.0

10.0

9.0

8.0

7.0

6.0

5.0

4.0100 200 550450350250150 500400300 600 kg

200 400 1,2001,000800600 lbm

S1

S2

Does Not Include Random Vibe

S3

PEG014

Acc

eler

atio

n (G

's)

Figure 4-5. Pegasus Net C.G. Load Factor Predictions.PEG027

10

Payload C.G. (Inches from Top of Payload Interface)

8

7

6

5

4

3

235252015 30 40

23" Sep System38" Sep System



payload if required. The nominal payload thermaland humidity environments for vehicle assembly,flight line, and captive carry operations are listedin Figure 4-10.

The component that exhibits the maximumtemperature inside the payload fairing, with aview factor to the payload, is the inner surface ofthe fairing. The temperature of the fairingincreases due to aerodynamic heating. Figure 4-11 shows the worst case transient temperatureprofile of the inner fairing surface adjacent to thepayload. The temperature profile was derivedusing the worst case heating trajectory, theminimum tolerance TPS thickness, and worstcase warm initial temperatures.

The component with a view factor to the payload,that exhibits the minimum temperature insidethe payload fairing, is also the inner surface of the

fairing. During captive carry, the payloadtemperature is primarily driven by radiativecooling. The fairing surface adjacent to thepayload can reach a minimum temperature of -40°C (-40°F). This temperature is reachedapproximately 30 minutes after OCA takeoff.

Fairing thermal emissivity on the inner surfacewill not exceed 0.9. As a non-standard service,a low emissivity coating can be applied to reduceemissivity to less than 0.1.

4.7.1 Nitrogen Purge

If required for spot cooling of a payloadcomponent, Orbital will provide localized GN2.The GN2 will meet Grade B specifications, asdefined in MIL-P-27401C and can be regulatedbetween 2.4-11.8 l/sec (5-25 scfm). The GN2 ison/off controllable at the LPO station. Onecooling location on the payload can be providedup to a total of 91 kg (200 lbm) of GN2 during taxiand captive carry. This cooling will be availablefrom payload mate through launch.

The system uses a ground nitrogen source until

Figure 4-6. Drop Transient Design Limit Load Environment.

Payload Interface ±0.5 g

Ax (g’s)Location

±0.5 g

Ay (g’s) Az (g’s)

±3.85

PE

G0

15

Figure 4-7. Payload Interface Random Vibration Specification.PEG016

0.1

0.01

0.001

0.0001

PowerSpectralDensity(g2/Hz)

Frequency (Hz)

10 100 1000 10000

0.0002

0.0005

0.002

0.005

0.02

0.05

X and Z Axes

Y Axis Only

20, .004 35, .004

40, .016 55, .016

1500, .004

2000, .001

65, .004

Frequency

(Hz)

20

35

40

55

65

1,500

2,000

Power Spectral

Density (g2/Hz)

X & Z Axes

Power Spectral

Density (g2/Hz)

Y Axis

.004

.004

.004

.004

.004

.004

.001

2.64

.004

.004

.016

.016

.004

.004

.001

2.69Overall Levels grms grms



OCA engine 2 starts, then it transfers to the OCAnitrogen system for captive carry. The system'sregulators are set to a desired flow rate, normally0.7 kg/min (1.5 lbm/min), then lockwired inplace. The system cannot be adjusted in-flight.This should be considered during payloadrequirement definition (i.e., volumetric flow ratewill increase as the OCA climbs to launchaltitude).

Payload purge requirements must be coordinatedwith Orbital via the ICD to ensure that therequirement can be achieved. Any payloadpurge requirement that cannot be met with theexisting system will be considered "out of scope"from the nominal Pegasus launch services.

4.8 Payload Electromagnetic Environment

All power, control and signal lines inside thepayload fairing are shielded and properlyterminated to minimize the potential for EMI.

The Pegasus payload fairing is radio frequency(RF) opaque, which shields the payload fromexternal RF signals while the payload isencapsulated. Based on analysis and supported

by test, the fairing provides 20 db attenuationbetween 1 and 10000 MHz. Figure 4-12 lists thefrequencies and maximum radiated signal levelsfrom vehicle antennas that are located near thepayload during powered flight. Antenna locatedinside the fairing are inactive until after fairingdeployment. Figure 4-13 lists carrier aircraftemitters and receivers. The payloadelectromagnetic environment (EME) results fromthree categories of emitters: Pegasus onboardantennas, Carrier Aircraft antennas, and Rangeradar. EME varies with mission phase. Forexample, the VAB environment is more benignthan the flight line/Carrier Aircraft environment.A worst case composite EME is defined in Figure4-14 and Figure 4-15, taking into account allmission phases. This EME should be comparedto the payload's RF susceptibility levels (MIL-STD-461, RS03) to define margin.

4.9 Payload Contamination Control

Orbital operates the Pegasus launch vehiclesystem under one of two contamination controlplans, depending on specific missionrequirements. These plans are:

Figure 4-8. Shock at the Base of the Payload.

100 200 300 500 1,000 2,000 3,000 5,000 10,00010

20

50

100

200

500

1000

2000

5,000

10,000

Frequency (Hz)

G’s

(10000, 3500)

(1000, 3500)

(100, 55)

Separating and Non-Separating Shock

PEG018



TD-0198 - “Pegasus Payload ContaminationControl Plan, Class 100,000 (Class M 6.5)Missions;” and

TD-289 - “Pegasus Payload ContaminationControl Plan, Class 10,000 (Class M 5.5)Missions.”

Class 10,000 (M 5.5) contamination control isavailable as a non-standard service.

These two plans are based on industry standardcontamination reference documents, includingthe following:

MIL-STD-1246C, “Product Cleanliness Levelsand Contamination Control Program”

FED-STD-209E, “Airborne Particulate CleanlinessClasses in Cleanrooms and Clean Zones.”

NRP-1124, “Outgassing Data for SelectingSpacecraft Materials”

The Pegasus vehicle and all payload integrationprocedures have been designed to minimize thepayload's exposure to contamination from thetime the payload arrives at the field integrationfacility through orbit insertion and separation.The VAB is maintained at all times as a visiblyclean, air-conditioned, humidity-controlled workarea.

As a nonstandard service, the payload can beprovided with a soft-walled cleanroom (SWC)with a Class 100,000 (Class M6.5) environmentfor payload integration operations at the VAB.Air is supplied to the SWC through a bank ofHigh-Efficiency Particulate Air (HEPA) filters,which are 99.97% effective in removing particlesof ≥0.3 microns in size. These filters are locatedin the ceiling of the enclosure from which air isdrawn from the VAB interior. Particulate size vs.time data is recorded in accordance with theguidelines of FED-STD-209E. The SWC is certified

12.5 20 80

90

100

110

120

130

Frequency (Hz)

Sou

nd P

ress

ure

Leve

l (db

)

OASPL = Overall Sound Pressure Level

Pegasus Carrier Aircraft Limit Envelope (OASPL = 124.8 dB)

Limit Level + 6 db(OASPL = 130.8 dB)

Figure 4-9. Payload Acoustic Environment.

114

111 109111.5

114116 117.5

121

125

120

117.5 117114

112.5

108

112

16 2531.5

4050

63 100125

160200

250315

400500

630800

1K1250

1.6K2K

2.5K3150

4K5K

PEG019



Pre-Payload Fairing Installation• Outside VAB Clean Tent• Inside VAB Clean TentPost-Payload Fairing Installation (GSE)• VAB• Roll-Out to Carrier Aircraft (VAFB)

• Carrier Aircraft Mate/Hot PadOCA AACS (Ground)• TaxiOCA AACS (Altitude)• Captive Carry• Abort/Contingency – Above 4.9 km (16 K ft) – Below 4.9 km (16 K ft)

23 ± 523 ± 5PLF Inlet23 ± 5Ambient(Generally 13 ± 11)23 ± 5PLF Inlet23 ± 5PLF Inlet23 ± 523 ± 5GN2

74 ± 1074 ± 10PLF Inlet74 ± 10Ambient(Generally 55 ± 20)74 ± 10PLF Inlet74 ± 10PLF Inlet74 ± 1074 ± 10GN2

A/CFiltered A/C

Filtered A/CFiltered AmbientFiltered A/C

Filtered A/CFiltered A/CFiltered A/CFiltered A/CFiltered A/CClean Nitrogen

45 ± 1545 ± 15

45 ± 15 <60(Note 1)

<70(Note 2)<50

None100 K (M6.5)

100 K (M6.5)

100 K (M6.5)

100 K (M6.5)

Purity Class(Note 3)

Humidity (%)ControlDeg FDeg C

Temp RangeEvent

Figure 4-10. Payload Thermal and Humidity Environment.

GSE A/C Performance Is Dependent Upon Ambient Conditions. Temperature Is Selectable and Controlled to Within ±2°C(±4°F) of Set Point. Resultant Relative Humidity Is Maintained to 45 ± 15%.AACS Ground Performance Is Dependent Upon Ambient Conditions (Dew Point). Temperature Is Selectable and ControlledWithin ±2°C (±4°F) of Set Point. Resultant Relative Humidity Is Maintained to 45 ± 15%.Class 10K (M5.5) Can Be Provided Inside the VAB Clean Tent and at the Payload Fairing Air-Conditioning Inlet on a MissionSpecific Basis As a Non-Standard Service.

1.

2.

3.

Notes:

PEG020

between 5 and 30 days prior to payload arrival atthe VAB.

During encapsulation, the payload fairing will beprovided with Class 100,000 air supplied by the

Figure 4-11. Pegasus XL Predicted Worst-Case Payload Fairing Inner Surface TemperaturesDuring Ascent to Orbit.

¥ Data Analytically Derived

¥ Worst Case Heating Profile (Hot Trajectory)

¥ Fairing Inner Surface Temperature at the Ogive/Cylinder Interface

0

Flight Time (Sec)

Temperature

(¡C)

Temperature

(¡F)

180

140

100200

250

300

350

100

150

50

0

60

20

-2025 50 75 100 125 150 175 200

PEG021



Figure 4-12. Pegasus XL RF Emitters and Receivers.

Function

RoleBandFrequency (MHz)

BandwidthPower Output SensitivityModulation

SourceCommand DestructReceiveUHF416.5 or425N/AN/A-107 dBmTone

TrackingTransponderTransmitC-Band5765

N/A400 W PeakN/APulse Code

TrackingTransponderReceiveC-Band5690

14 MHz at 3 dBN/A-70 dBmPulse Code

InstrumentTelemetryTransmitS-Band2269.5

750 KHz at 3 dB5 WN/AFM/FM

BoosterTelemetryTransmitS-Band2288.5

315 KHz at 3 dB5 WN/APCM/FM

GPS

ReceiveL-Band1575.42 – 10.231227.60 – 10.2320.46 MHzN/AN/APRN Code

Camera

TransmitS-Band2200-2400

12 MHz8 WN/AN/A

1 2 3 4 5 6 7

PEG022

VAB air conditioning HEPA system. A diffuser isused at the fairing inlet to direct the airflow awayfrom the payload. During Pegasus transport tothe OCA and during Pegasus/OCA mate, a blower/desiccant system provides Class 100,000 air tothe fairing. These blowers process ambient airthough a desiccant canister and a HEPA filter.For hot pad operations after Pegasus/OCA mate,the Ground Air Conditioning System (GACS) isused; during taxi and captive carry on the OCA,the aircraft's Airborne Air Conditioning System(AACS) is used. Both deliver HEPA-filtered Class100,000 air to the fairing, and both employ adiffuser to direct the airflow away from thepayload. The face velocity will not exceed 11 m/min (35 ft/min).

Particle count measurements will be made foreach fairing air supply (i.e. - the VAB air supply,the blower/desiccant system, the GACS, and the

AACS) before hookup to the fairing. Thiscertification will be made after each system hasbeen running a minimum of 30 minutes, toensure that the downstream ducting has beenpurged.

Also as a non-standard service, carbon filters canbe provided to remove volatile hydrocarbons ofmolecular weight 70 or greater from the fairingair supply, with better than 95% efficiency.

The Pegasus payload fairing inner surface isconstructed of graphite/epoxy compositematerial, meeting the NRP-1124 outgassingstandards of Total Mass Loss (TML) ≤1.0%, andCollected Volatile Condensable Material (CVCM)≤ 0.1%.

The baseline cleanliness of the fairing innersurface is “visibly clean.” “Visibly clean” isdefined as appearing clean of all particulate and

Figure 4-13. Carrier Aircraft RF Emitters and Receivers.

Function

Role

Band

Frequency

(MHz)

Bandwidth

Power Output

Sensitivity

Modulation

Source

HF Comm

Receive/

Transmit

HF

2.0-29.999

SSB: 3 KHz

AM: 6 KHz

SSB

AM

SSB: 1mV

AM: 3mV

SSB: 400 W

AM: 125 W

UHF Comm

Receive/

Transmit

UHF

225.0-399.975

Standard A/C

Radio

10 W

4 mV

AM

GPS/Loran

Receive

L-Band

1575.42

20.46 MHz

N/A

N/A

PRN Code

VHF Comm

Receive/

Transmit

VHF

118.0-135.975

90 KHz

@-60 dB

25 W

3mV

AM

ATC

Transponder

Receive/

Transmit

L-Band

R: 1030 ±0.2

T: 1090 ±3

25 MHz @ -60 dB

500 W

-76 dBm

Pulsed 1% Duty

Cycle

Video

Telemetry

Transmit

S-Band

2210.50

or 2383.5

12 MHz

10 W

N/A

FM

GPS

Relay

Receive/

Transmit

L-Band

1575.42

20.46 MHz

<1 W

N/A

PRN Code

1 2 3 4 5 76 8

Weather

Radar

Receive/

Transmit

X-Band

9345 ±30

R: 700 KHz

65 kW

Not

Specified

5.74 mS Pulse,

200 pps

PEG023



100

10

1

0.1

0.010.01 0.1 1 10 100 1000 10000

Frequency (MHz)

Fie

ld (

V/m

)

Figure 4-14. Western Range Worst Case Composite Electromagnetic Environment.PEG024

nonparticulate substances when examined bynormal 20/20 vision at a distance of 15-46 cm (6-

18 in) under incident light of 1,076-1,346 lux(100-125 foot-candles).

Level 750A, Level 600A, or Level 500A cleanlinessrequirements of MIL-STD-1246C can be providedas a non-standard service.

4.10 Payload Deployment

Following orbit insertion, the Pegasus avionicssubsystem can execute a series of pre-programmed Reaction Control System (RCS)commands from the MDL to provide the desiredinitial payload attitude prior to payload separation.This capability may also be used to incrementallyreorient for the deployment of multiple spacecraftwith independent attitude requirements. Eitheran inertially-fixed or spin-stabilized attitude maybe specified by the user.

Pegasus can accommodate a variety of payloadspinup requirements up to 60 rpm. The maximumrate for a specific mission depends upon the spinaxis moment of inertia of the payload and theamount of nitrogen needed for other attitudemaneuvers. Figure 4-17 shows the accuracy ofcontrol and spin rate. Post-separation rate errors

Source CommentFrequency

(MHz)Field(V/m)

Figure 4-15. Worst Case Composite Electromagne-tic Environment.

HF Comm

VHF Comm

ATC Transponder

ARSR-1E

PEG S-Band

AN/GPN-12

Range C-Band (Tracking Transponder)

PEG C-Band

Range C-Band (Skin Tracking)

Weather Radar

Aircraft Communications

Aircraft Communications

Aircraft Transponder (Pathfinder Data)

VAFB Air Surveillance Radar

Vehicle Accel/Telemetry Transmitter

VAFB RAPCON

FPS-16-1, TPQ-18, HAIR, MOTR, FPQ-6, MPS-36 (Pathfinder Data)

Vehicle Transponder

Multiple Objects Tracking Radar (Pathfinder Data)

OCA During Pathfinder, Full Slew

2.8-26

118-136

1090

1320

2269.5, 2288.5

2800

5690

5765

5890

9345

<0.16

<0.1

1.7

0.4

9.0

8.8

8.7

28.0

2.3

<6

PE

G02

5



Error Type (Pegasus

Vehicle Axes)

Figure 4-16. Typical Pre-Separation PayloadPointing and Spin Rate Accuracy.

Angle

(Degrees)

Rate

(Degrees

per Sec)

Pointing

±4

±4

±4

Ð

Yaw (Z)

Pitch (Y)

Roll (X)

Spin Rate

±0.5

±0.5

±1.0

±2.0

Notes:

(1)

(2)

Accuracies are Dependent on Payload Mass Properties.

Pointing Angle of ±4° Is for Sun-Pointing Payloads. ForNon-Sun-Pointing Payloads, Accuracies of ±3° ArePossible. PEG026

are dependent on payload mass properties.

4.11 Payload Tip-off

Payload tip-off refers to the angular velocityimparted to the payload upon separation due toan uneven distribution of torques and forces.

If a Marmon Clamp-band separation system isused, payload tip-off rates are generally under 4°/sec per axis. This can vary depending on themass properties of the payload and theconfiguration of the separation system. Orbitalperforms a mission-specific tip-off analysis foreach payload.

Section 5.0—Spacecraft Interfaces


Pegasus User's Guide Section 5.0—Spacecraft Interfaces

5.1 Payload Fairing

This section describes the fairing, fairingseparation sequence, payload dynamic envelope,and payload access panel. The standard payloadfairing consists of two graphite composite halves,with a nosecap bonded to one of the halves, anda separation system. Each composite half iscomposed of a cylinder and an ogive section.The two halves are held together by two titaniumstraps, both of which wrap around the cylindersection, one near its midpoint and one just aft ofthe ogive section. Additionally, an internalretention bolt secures the two fairing halvestogether at the surface where the nosecap overlapsthe top surface of the other fairing half. The baseof the fairing is separated using a non-contaminating frangible joint. Severing thealuminum attach joint allows each half of thefairing to then rotate on hinges mounted on theStage 2 side of the interface.

5.1.1 Fairing Separation Sequence

The fairing separation sequence consists ofsequentially actuating pyrotechnic devices thatrelease the right and left halves of the fairing froma closed position, and deploy the halves awayfrom either side of the core vehicle. The nose boltis a non-contaminating device. The pyrotechnicdevices include a separation nut at the nose,forward and aft bolt cutter pairs for the externalseparation straps at the cylindrical portion of thefairing, a frangible joint separation system at thebase, and a pyrogen gas thruster system fordeployment.

5.1.2 Payload Dynamic Design Envelope

The fairing drawings in Figures 5-1 and Figures5-2 show the maximum dynamic envelopesavailable for the payload during captive-carryand powered flight for the XL and HAPSconfigurations. The dynamic envelopes shownaccount for fairing and Pegasus structuraldeflections only. The customer must take intoaccount payload deflections due tomanufacturing/design and tolerance stack-upwithin the dynamic envelope. Proposed payloadenvelope violations must be approved by Orbital.

No part of the payload may extend aft of thepayload interface plane without specific Orbitalapproval. These areas are considered stayoutzones for the payload and are shown in Figure 5-1 and Figure 5-2.

Incursions to these zones may be approved on acase-by-case basis. Additional analysis is requiredto verify that the incursions do not cause anydetrimental effects. Vertices for payload deflectionmust be given with the Finite Element Model toevaluate payload dynamic deflection with theCoupled Loads Analysis (CLA). The payloadcontractor should assume that the interface planeis rigid; Orbital has accounted for deflections ofthe interface plane. The CLA will verify that thepayload does not violate the dynamic envelope.

5.1.3 Payload Access Door

Orbital provides one 21.6 cm x 33.0 cm (8.5 in x13.0 in), graphite, RF-opaque payload fairingaccess door. The door can be positioned accordingto user requirements within the zone defined inFigure 5-3. The position of the payload fairingaccess door must be defined no later than L - 8months.

5.2 Payload Mechanical Interface and SeparationSystem

Orbital will provide all hardware and integrationservices necessary to attach non-separating andseparating payloads to Pegasus. All attachmenthardware, whether Orbital or customer provided,must contain locking features consisting of lockingnuts, inserts or fasteners. Orbital provides identicalbolt patterns for both separating and non-separating mechanical interfaces.

5.2.1 Standard Non-Separating MechanicalInterface

Figure 5-4 illustrates the standard, non-separatingpayload mechanical interface. This is for payloadsthat provide their own separation system andpayloads that will not separate. Direct attachmentof the payload is made on the Avionics Structurewith sixty 0.48 cm (0.19 in) fasteners as shown inFigure 5-4. Orbital will provide a matched drilltemplate to the payload contractor to allow



212.983.8

110.043.3

10.04.0

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

Stayout ZoneClamp/SeparationSystem Components

Payload Interface Connector

f 97 cm (38 Inches) Payload Separation System

Stayout Zone

HarnessPigtails

to Payload

0°

270°90°

180°

Payload Interface Planefor Non-Separating Payloads

Payload Interface Planefor Payload Separation

System

97 cm (38 in) Avionics Structure56 cm (22 in) Long

116.846.0

RCS StayoutZone

Pyrotechnic Event Connector

77.730.6

+X

+Z

Side View

Forward ViewLooking Aft

Notes: (1) Fairing Door Location Is Flexible Within a Specific Region. (Figure 5-3).

(2) Payload Must Request Any Envelope Below Bolted Interface.

(3) If Payload Falls within RCS Controllability Dead Band They Must Honor RCS Stayout Zone.

(4) If the Payload Requires Nitrogen Cooling, then the Payload Envelope Will be Locally Reduced by 1 Inch Along the Cooling Tube Routing.

PayloadStayout Zones

Legend:

R 269.2 106.0

PayloadDynamicEnvelope

Fairing

100.3 39.5

f

f

f

76.029.9

2.51.0

Dimensions in cmin

Ogive MateLine

PEG028



Payload Interface Planefor Non-Separating Payloads

Payload Interface Planefor Payload Separation System

97 cm (38 in) Avionics Structure

72.4 cm (28.5 in) Long

Fairing

76.029.9

R 269.2106.0

Figure 5-2. Payload Fairing Dynamic Envelope with Optional Hydrazine Auxiliary Propulsion System(HAPS) and 97 cm (38 in) Diameter Payload Interface.

186.273.3

(S3/HAPS Sep Joint Included)

(S3/HAPSSep Joint Inc.)

10.04.0

Ogive Mate

Line

PayloadDynamicEnvelope

Dimensions in

Side View

+X

+Z

Notes: (1) Fairing Door Location is Flexible Within a Specific Region. (See Figure 5-3)

(2) Payload Must Request Any Envelope Below Bolted Interface.

83.332.8

f 100.3 39.5

f 116.8 46.0

cmin

PEG029



Figure 5-3. Payload Fairing Access Door Placement Zone.

Pegasus

Station X

+1,573.4

+619.4

Pegasus Access

Door Zone

Pegasus

Station X

+1,493.7

+588.1

13

5

13

5Arc Length

ArcLength

Applies on Either

Side of Fairing Joints at

0° and 180°

Dimensions incm

in

+X

+Z

Pegasus CoordinatesNotes:38" Payload

Interface Place

Pegasus Station X

(cm/in)

23" Payload

Interface Place

Pegasus Station X

(cm/in)

(1) Entire Access Hole Must Be Within Specified Range.

(2) One 21.6 cm x 33.0 cm (8.5 in x 13.0 in) Door perMission Is Standard.

(3) Edge of Door Cannot Be Within 13 cm (5 in) ofFairing Joints

Separable

Non-Separable

1485.4/584.8

1475.4/580.9

1509.4/594.3

1501.9/591.3

PEG030



Figure 5-4. Non-Separable Payload Mechanical Interface.

Payload Harness

Pegasus

Stage 3

Harness

Forward

+X

Harness Access Hole

Forward Interface of f 97 cm (38 in),

56 cm (22 in) Long Avionics Structure

Rotated 135° CWApplies at 45° (Pyrotechnic Event) and

225° (Payload Interface)

22.9

9.0

45°

98.6

38.8

270°

225°

180°

90°

116.8

46.0

Fairing Dynamic Envelope

MS27474T-16F-42S

(Payload Interface Connector)

Payload

Stayout

Zone

MS27474T-14F-18S

(Pyrotechnic Event

Connector)

Bolt Circle Consists of

60 0.51 cm (0.20 in)

Holes Equally Spaced,

Starting at 0°

45°

0°

+Y

+Z

10.3

4.1

Pegasus

Coordinates

Forward View

Looking Aft

f

f

Dimensions incm

in

5.7 ± .09

2.3 ± .04f

1.90

21.25

2X19.95

PEG031



accurate machining of the fastener holes and willsupply all necessary attachment hardware perthe payload specifications. The Orbital provideddrill template is the only approved fixture fordrilling the interface. The payload contractorwill need to send a contracts letter requestinguse, on a non-interference basis, of the drilltemplate (no later than 30 days prior to neededdate). The payload contractor should plan ondrill template usage for a maximum of two weeks.

5.2.2 Standard Separating Mechanical Interface

If the standard Pegasus payload separation systemis used, Orbital controls the entire spacecraftseparation process. The standard separationsystem uses a Marmon clamp design. Threedifferent separation systems are available,depending on payload interface and size. Theyare 97 cm (38 in), 59 cm (23 in), and 43 cm (17in) separation systems. The 97 cm (38 in)separable payload interface is shown in Figure 5-5; the 59 cm (23 in) separable payload interfaceis shown in Figure 5-6; the 43 cm (17 in) separablepayload interface is shown in Figure 5-7.

The separation ring to which the payload attachesis supplied with through holes. The weight ofhardware separated with the payload isapproximately 4.0 kg (8.7 lbm) for the 97 cm (38in) system, 2.7 kg (6.0 lbm) for the 59 cm (23 in)system, and 2.1 kg (4.7 lbm) for the 43 cm (17 in)system. Orbital-provided attachment bolts tothis interface can be inserted from either thelaunch vehicle or the payload side of this interface(NAS6303U, dash number based on payloadflange thickness). The weight of the bolts, nuts,and washers connecting the separation system tothe payload is allocated to the separation system.Orbital will provide a matched drill template tothe payload contractor to allow accuratemachining of the fastener holes and will supplythe integration ring and all necessary attachmenthardware to payload specifications. The payloadcontractor will need to send a contracts letterrequesting use, on a non-interference basis, ofthe drill template (no later than 30 days prior toneeded date). The payload contractor shouldplan on drill template usage for a maximum of

two weeks. The flight separation system shall bemated to the spacecraft during processing at theVAB.

At the time of separation, the flight computersends commands which activate redundant boltcutters, which allows the titanium clampbandand its aluminum shoes to release. The band andclamp shoes remain attached to the avionicsstructure by retention springs. The payload isthen ejected by matched push-off springs withsufficient energy to produce the relative separationvelocities shown in Figure 5-8. If non-standardseparation velocities are needed, different springsmay be substituted on a mission-specific basis.

5.3 Payload Electrical Interfaces

5.3.1 Umbilical Interfaces

A block diagram of the standard Pegasus electricalinterface capabilities is shown in Figure 5-9.

The standard payload electrical connector andharness configuration is shown in Figure 5-10and in Figure 5-11. Note that two interfaceconnectors are used to implement the standardinterface. The formal electrical interface is definedas the separation plane of the connectors. Orbitalwill provide the payload side of the interfaceconnectors (payload side - MS27474T-16F-42Sfor telemetry, and MS27474T-14F-18S forpyrotechnic commands) one year prior to launch.The payload should integrate these connectors tothe spacecraft flight harness forward of theinterface plane. This harness will be integrated tothe separation system by Orbital two monthsbefore launch. The matching interface connectorsand the associated electrical harnesses aft of theinterface plane are provided by Orbital for non-separating and separating payloads.

All interface wires are shielded for EMI protection.The Orbital flight harnesses and payload-providedharness will be integrated with the flight separationsystem and will be available no earlier than onemonth prior to launch. The separation system/harnessing will be mated to the spacecraft at theVAB. Physical and functional testing of theharnessing will be accomplished on a mutually



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

Dimensions in cmin

Maximum Allowable Payload = 454 kg (1,000 lbs) (Shear Critical)

Forward View Looking Aft

PayloadPush-OffSprings (4 Places)

Clamp BandRetention Springs (8)

Bolt Cutters (2)(Redundant)

Payload Interface

270°

180°

0°

+Z

+Y

PegasusCoordinates

18 Pin Payload Pyro Connector

(MS-27474-14F-18S)

42 Pin Payload Umbilical Connector (MS-27474-16F-42S)

Bolt Circle Consists of60 0.48 cm (0.19 in) Holes Equally Spaced, Starting at 0°

98.58

38.81

10.0

4.05.0

2.0

Avionics Structure

Payload Separation

Clamp Band

Payload Interface

PlaneSeparation

Plane

+Y

+X

4.0 kg (8.7 lbm) Remains with

Payload (Includes Harness)

Side View

f Bolt Circle

PEG032

90°




180°

0°Payload Push-Off

Springs (4 Places)

Retention Springs (8)

Payload Interface

Clamp Band

Bolt Cutters (2)

(Redundant)

Adpater Cone

42 Pin Payload

Umbilical Connector

(MS-27474-16F-42S)

90°

Maximum Allowable Payload = 317 kg (700 lbs) (Shear Critical)


18 Pin Payload Pyro Connector

(MS-27474-14F-18S)


32 0.64 cm (0.25 in)


Starting at 0°

+Z

+Y

Pegasus

Coordinates

270°

Side View

Adapter Cone

Bolt Cutters (2)

(Redundant)


Payload Separation Clamp Band

Payload Attachment

Plane

Bolt Circle

3.75

1.48

Separation

Plane

2.7 Kg (6.00 lbm)

Remains with Payload

(Includes Harness)

7.49

2.95

59.06

23.25f

Dimensions incm

in

+X

+YPEG033




180°

270°

Clamp Band

Payload Push-Off

Springs (12 Places)

0°

Adapter

Cone

Bolt Cutters (2)

(Redundant)

90°



Maximum Allowable Payload = 193 kg (425 lbm) (Shear Critical)


24 0.64 cm (0.25 in)


Starting at 0¡

Dimensions incm

in

Payload Attachment Plane

Side View from 0¡

Separation

Plane

Bolt Circle

3.75

1.48

Adapter Cone

Payload Separation

Clamp Band

Retention Springs

2.1 Kg (4.7 lbm)

Remains with Payload

(Includes Harness)

Bolt Cutters (2)

(Redundant)

7.49

2.95

43.2

17.0f

99.1

39.0f

Pegasus

Coordinates

+X

+YPEG034



agreed to schedule between Orbital and thepayload customer.

5.3.2 Payload Auxiliary Power

Payloads can receive power during groundoperations and captive flight directly from thecarrier aircraft or from payload-provided AirborneSupport Equipment (ASE) via the five standardtwisted shielded pair pass-throughs mounted tothe Stage 3 avionics structure. The lines interfacewith the Pegasus or payload ASE through thePegasus wing interface. The payload isaccountable for the weight of cables on thepayload side of the interface and all non-standardcables. No power is available during transport.

5.3.3 Payload Command and Control

Discrete sequencing commands, generated bythe Pegasus flight computer, are available to thepayload. These commands are opto-isolatedpulses of programmable lengths in multiples of40 ms. Up to eight command line pairs, eachcapable of multiple pulses, can be provided forthe payload. Discrete lines are provided throughthe same interface connector used for the payloadauxiliary power lines (connector MS27484T-16F-42P). The payload supplies the voltage (≤40

VDC) and must limit current to 500 milliamps(mA) nominal in a fashion similar to using a drycontact relay.

5.3.4 Payload Status Monitoring

Payload discrete telemetry downlink can beprovided during ground processing (limited toPegasus on-times), checkout, captive carry andduring launch as part of the standard service. Upto four discrete telemetry signals can beaccommodated in the Pegasus telemetry streamvia the flight computer. This telemetry interfaceincludes a signal and ground for each discretetransmitted on dedicated twisted shielded wirepairs. The flight computer contains a resistor tolimit the current of a 5.0 VDC signal toapproximately 10 mA. The interface must beoptically isolated at the payload. See Section 10for a description of the serial telemetry linkoption.

5.3.5 Payload Pyrotechnic Initiator Driver Unit

For a standard mission, one dual and four single75 ms pulses at 5 amps are available for post-launch use by the spacecraft. Use of the standardseparation system requires two of the singleoutputs. The firing commands are sent via the

PEG035

Payload Weight

.5

.75

1.00

1.25

1.5

1.75

Figure 5-8. Payload Separation Velocities Using the Standard Separation System.

Sep

arat

ion

Vel

ocity

(m

/sec

)

Sep

arat

ion

Vel

ocity

(ft/

sec)

600

2.00

3.00

4.00

5.00

1,2001,0008006004002000 lbm

kg5004003002001000

97 cm (38 in) Interface





Pegasus avionics subsystem Pyro Driver Unit(PDU). The pyro interface is provided through aseparate connector from the power/commandconnector.

5.3.6 Range Safety Interfaces/Vehicle FlightTermination

The Pegasus air-launched approach minimizesinterfaces with the test range. All ordnance onthe Pegasus vehicle is in the safe condition whilein captive carry mode under the carrier aircraft.Ordnance is armed during a sequence which isinitiated upon release from the OCA. Proceduresfor arming ordnance on the spacecraft aredetermined on a mission-specific basis. Noarming of the payload prior to drop from thePegasus Carrier Aircraft is allowed.

Generally, the standard Pegasus FTS subsystem

satisfies all range safety requirements withoutadditional FTS support from the payload.However, information on the payload, such as abrief description, final orbit, spacecraft ordnance,hazardous operations and materials summary,will be requied to support range documentation.Additional range support for payload operations,such as orbit determination and command andcontrol, can be arranged. Range-providedservices have long lead times due to Departmentof Defense (DoD) and NASA supportrequirements; therefore, test range supportrequirements must be identified early in order forOrbital to ensure their availability.

5.3.7 Electrical Power

Power lines shall be isolated from the Pegasus XLand payload structures by at least 1 megohm.

• 5 DPDT Switches• 0 - 55 ±5 VDC < 18 Amps

OrbitalCarrier

Aircraft LPO

Opto-Isolation

5 Twisted Shielded Pairs (22 AWG)

Flight Computer

Pegasus Payload

Control

Status

VCC

8 Standard

4 Standard

PDU5 VDC, 5A, 75 msec, 6 Pyro Events

(1 Dual Output + 4 Single Outputs)

Output +

Output -

Sep Loop

Sep Loop

42 P

in C

onne

ctor

Power/Signal• Pass-Throughs• Captive Carry• < 3A/Pair

Discrete Commands• Max Voltage Switching 45 VDC• Max Transient Voltage 60 VDC• Max Current Switching 0.5 ADC• Max Turn-on Time 1.5 ms• Max Turn-off Time 0.25 ms• Max Leakage 40 µA• High and Low Side Switching• Short Circuit Protected

Talkbacks• Continuity or Switch On: < 0.5 VDC at 10 mA• Open or Switch Off: High Impedance, > 100 K

Figure 5-9. Pegasus Payload Electrical Interface.

18 P

in

PEG036



Plug Shell

Receptacle Shell

Socket Contacts

Pin Contacts

SP

S P P S

S P

Separation Plane

Mate #1 Performedat Orbital During

Separation System Assembly

Mate #2Performed

at VAB

Plug withPin Contacts

MS-27484T-16F-42P

Receptacle withSocket Contacts

MS-27474T-16F-42S

PayloadInterface

Plane

Spacecraft

HarnessLengthSpecified byPayload

Plug withSocket Contacts

MS-27484T-16F-42S

Receptacle withPin Contacts

MS-27474T-16F-42P

Supplied toPayload

(RecommendHard Mount)

Figure 5.10. Pegasus/Spacecraft Electrical Connectors and Associated Electrical Harnesses.

Launch Vehicle

Note: Sep System and Pigtails Delivered to VAB as a Unit

PEG037

Plug Shell

Receptacle Shell

Socket Contacts

Pin Contacts

SP

S P P S

S P

Separation Plane

Mate #1 Performedat Orbital During

Separation SystemAssembly

Mate #2Performed

at VAB

Plug withPin Contacts

MS-27484T-14F-18P

Receptacle withSocket Contacts

MS-27474T-14F-18S

PayloadInterface

Plane

Spacecraft

HarnessLengthSpecified byPayload

Plug withSocket Contacts

MS-27484T-14F-18S

Receptacle withPin Contacts

MS-27474T-14F-18P

Supplied toPayload

(RecommendHard Mount)

Figure 5.11. Pegasus/Spacecraft Pyrotechnic Connectors and Associated Electrical Harnesses.

Launch Vehicle

PEG038



The Launch Vehicle System (the Pegasus XL, theintegration site facilities and the OCA) and SpaceVehicle System (the payload and all groundbased systems required to process, launch andmonitor the payload during all phases of launchprocessing and flight operations) shall each utilizeindependent power sources and distributionsystems.

5.3.8 Electrical Dead-Facing

Prior to T-0, all Space Vehicle System electricalground support equipment electrical interfacesat the umbilical shall be dead-faced to ensurethat there shall be no current flow greater than 10mA across the umbilical interface. Prior to drop,all aircraft power shall be isolated from thelaunch vehicle and the payload.

5.3.9 Pre-Separation Electrical Constraints

Prior to initiation of the separation event, allpayload and launch vehicle electrical interfacecircuits shall be constrained to ensure that thereshall be no current flow greater than 10 mA DCacross the separation plane during the separationevent.

5.3.10 Non-Standard Interfaces

Additional interface options are available. SeeSection 9.0 for a description.

5.4 Payload Design Constraints

5.4.1 Payload Center of Mass Constraints

To satisfy structural constraints on the standardStage 3 avionics structure, the axial location ofthe payload center of gravity (c.g.) along the Xaxis is restricted as shown in Figure 5-12. Alongthe Y and Z axes, the payload c.g. must be within3.8 cm (1.5 in) of the vehicle centerline for thestandard configuration and within 2.5 cm (1.0 in)of centerline if HAPS is used (including tolerancesin Figure 5-13). Payloads whose c.g. extend

Measurement

Figure 5-13. Payload Mass Property MeasurementError Tolerances.

Mass

Principal Moments of Inertia

Cross Products of Inertia

Center of Gravity X, Y and Z Axes

Error Tolerance

±0.5 kg (±1 lbs)

±5%

±0.7 kg - m2 (±0.5 sl - ft2)

±6.4 mm (±0.25 in)

PE

G04

2

0 25 50 75 100 125 cm

0 10 20 30 40 50 in

0 0

500

800

600

200

400

1,200

C.G. Location From Interface Plane

Pay

load

Mas

s (k

g)

Pay

load

Mas

s (lb

m)

200

100

300

400

Non-Separating

38"

38" With HAPS

23"

17"

1,000

Figure 5-12. Payload Mass vs. Axial C.G. Location on X Axis.PEG039A



beyond these lateral offset limits will requireOrbital to verify that structural and dynamiclimitations will not be exceeded. Payloads whoseX-axis c.g. falls into the RCS Dead Band Zonereferred to in Figure 5-14 will require movementof the RCS thrusters which can be supported ona mission-specific basis.

Mass property measurements must adhere to thetolerances set forth in Figure 5-13. The payloadcenter of mass must not transition through theRCS Dead Band Zone during the unpoweredflight (before stage ignition or after burnout) orloss of attitude control capability will occur.

5.4.2 Final Mass Properties Accuracy

The final mass properties statement shall specifypayload weight to an accuracy of 0.5 kg, thecenter of gravity to an accuracy to 6.4 mm ineach axis, and the products of inertia to 0.7 kg-m2. In addition, if the payload uses liquidpropellant, the slosh frequency must be providedto an accuracy of 0.2 Hz, along with a summaryof the method used to determine slosh frequency.

5.4.3 Payload EMI/EMC Constraints

The Pegasus avionics shares the payload areainside the fairing such that radiated emissionscompatibility is paramount. The Pegasus avionics

RF susceptibility levels have been characterizedby test. Orbital places no firm radiated emissionslimits on the payload other than the prohibitionagainst RF transmissions within the payloadfairing. Prior to launch, Orbital requires reviewof the payload radiated emission levels (MIL-STD-461, RE02) to verify overall launch vehicleEMI safety margin (emission) in accordance withMIL-E-6051. Payload RF transmissions are notpermitted after fairing mate and prior to separationof the payload. An EMI/EMC analysis will berequired to ensure RF compatibility.

Payload RF transmission frequencies must becoordinated with Orbital and range officials toensure non-interference with Pegasus and rangetransmissions. Additionally, the customer mustschedule all RF tests at the integration site withOrbital in order to obtain proper range clearancesand protection.

5.4.4 Payload Stiffness

To avoid dynamic coupling of the payload modeswith the 8-9 Hz natural frequency of the PegasusXL vehicle, the spacecraft should be designedwith a structural stiffness to ensure that thefundamental frequency of the spacecraft, fixed atthe spacecraft interface, in the Pegasus Z axis isgreater than 20 Hz.

Pay

load

Cen

ter

of M

ass

Offs

et

(Rel

ativ

e to

For

war

d In

terf

ace

of ø

38"

, 22"

Lo

ng A

vion

ics

Str

uctu

re)

250

200

150

cm in

100

50

100

90

60

40

20

Figure 5-14. Detailed RCS Dead Band Zone.

00 100

200 400 600 800 lbm

200 300 400 kg

Payload Weight

Pegasus RCS Stay-Out Zone Will Apply to Payloads Which Have a Center of Mass Offset in the Shaded Area

PEG040



5.4.5 Payload Propellant Slosh

A slosh model should be provided to Orbital ineither the pendulum or spring-mass format. Dataon first sloshing mode are required and data onhigher order modes are desirable.

5.4.6 Customer Separation System ShockConstraints

If the payload employs a non-Orbital separationsystem, then the shock delivered to the PegasusStage 3 vehicle interface must not exceed thelimit level characterized in Figure 4-3. Shockabove this level could require a requalification ofunits or an acceptance of risk by the payloadcustomer.

5.4.7 System Safety Constraints

Orbital considers the safety of personnel andequipment to be of paramount importance. Thepayload organization is required to conduct atleast one dedicated payload safety review inaddition to submitting to Orbital an AccidentRisk Assessment Report (ARAR) or equivalent asdefined in EWR 127-1.

Organizations designing payloads that employhazardous subsystems are advised to contactOrbital early in the design process to verifycompliance with system safety standards.

EWR 127-1 and WFF RSM-93 outline the safetydesign criteria for spacecraft on Pegasus vehicles.These are compliance documents and must bestrictly adhered to. It is the responsibility of thepayload contractor to insure that the payloadmeets all Orbital and range imposed safetystandards.

5.5 Carrier Aircraft Interfaces

5.5.1 Payload Services

The OCA can provide DC power to the payloadduring flight line operations and captive carry.This power is supplied by the OCA through thepayload interface connector mounted to the Stage3 avionics structure, as described in Section 5.3.Figure 5-14 provides details on the Pegasus/OCAinterface.

Orbital provides up to five twisted shielded pairsof pass-through wires (22 AWG) to the Launch

Launch PanelOperator Station

Air ConditioningSystem Pallet

AvionicsPallet

Wire HarnessUmbilicals

NitrogenPurge/

CoolingReservoir

AACSInlet

Payload Fairing

Nitrogen PurgeManifold

PegasusLaunchVehicle

CarrierAircraft

Pay-load LPO

Station

SeparationPlane

PegasusWing

5 Twisted PairPass Throughs

8 DiscreteCmds

4 Talkbacks

Pyro Events

Figure 5.14. Pegasus/OCA Interface Details.PEG110



Panel Operator (LPO) Station as a standard servicein the aircraft.

Orbital provides on-board payload monitoringcapabilities through the Orbital-manned LPOstation. The LPO station is equipped withcommunications and safety equipment, and canaccommodate flight qualified rack-mountedpayload support equipment if required.

5.5.2 Payload Support at Launch Panel OperatorStation

The Pegasus Launch Panel Operator (LPO) Stationprovides a 48 cm (19 in) rack for payload specificairborne support equipment (ASE), up to amaximum volume equivalent to two rack-mounted PCs. Payload ASE must comply withMIL-STD-810D. The payload rack is suppliedwith four 5A circuits of unregulated 28 VDCpower plus one 5A circuit of 115 VAC, 400 Hzpower. Additional equipment provided includesan adjustable DC power supply and a switchpanel. The power supply features a selectablevoltage level of 0-55 ±5 VDC and a 0 to 18Aadjustable current limit. Digital displays indicateboth voltage and current. Maximum allowablecurrent is limited to 3A per twisted, shielded pairof pass-through wires. The switch panel containstwelve double-pole, double-throw switches withfive amp contacts. Five of the switches havemomentary actuation. The seven remainingswitches have alternate actuation. The switchpanel is provided with two 5A circuits ofunregulated 28 VDC power. No provisions areavailable for seating a payload representative atthe LPO Station in-flight. The Pegasus LPO willbe available to perform limited payload operationsduring non-critical portions of the flight checklist,as defined in the Mission Integration WorkingGroups (MIWGs) and documented in the LPOChecklist.

Section 6.0—Mission Integration


Pegasus User's Guide Section 6.0—Mission Integration

6.1 Mission Management Structure

Successful integration of payload requirements isparamount in achieving complete missionsuccess. Pegasus has established a mission teamapproach to ensure all customer payloadrequirements and services are provided. As themission evolves the team is responsible fordocumenting, tracking and implementingcustomer requirements and changes. AConfiguration Control Board (CCB) ensures theserequirements are supportable and appropriatelyimplemented. The Pegasus mission team isresponsible for providing the customerrequirements, as well as changes to theserequirements, to the CCB. Open communicationbetween the Pegasus and payload customer isessential for ensuring total customer satisfaction.To facilitate the necessary communication andinteraction, the Pegasus mission integrationapproach includes establishing a mission team,holding technical meetings and supportingreadiness reviews.

An organizational structure is established foreach Pegasus mission to manage payloadintegration, mission preparations and executethe mission. Open communication betweenOrbital and the customer, emphasizing timelytransfer of data and prudent decision-making,ensures efficient launch vehicle/payloadintegration operations.

The Orbital and customer roles in missionintegration is illustrated in Figure 6-1. The ProgramManagers, one from the customer and one fromOrbital, execute the top-level management duties,providing overall management of the launchservices contract. Within each organization, oneperson will be identified as the Mission Managerand will serve as the single point of contact intheir respective organizations for that mission.The customer should appoint a Payload MissionManager within its organization. All payloadintegration activities will be coordinated andmonitored by the Mission Managers, includingmission planning, launch range coordination,

Figure 6-1. Mission Integration Management Structure.

PayloadProgram Manager

PayloadMission Manager

Payload ProgramTechnical Support

Pegasus SystemsEngineering

Electrical AnalysisSystems Integration

Mission Analysis

Mechanical Analysis

PegasusMission Engineer

Procedure Preparation

Mission Requirements

Production Planning

Mission Integration

Pegasus LaunchSite Operations

Facilities Management

Systems Testing

Vehicle Integration

Safety & QA

PegasusMission Manager

Launch Operations

Payload Requirements

Range Coordination

PegasusProgram Manager

PegasusContracts Manager

PegasusLaunch Services

Director

PayloadContracts Manager

Mission Interface

PEG043



and launch operations. The Payload MissionManager is responsible for identifying the payloadinterface requirements and relaying them to thePegasus Mission Manager. The Pegasus MissionManager is responsible for ensuring all the payloadlaunch service requirements are documentedand met. Supporting the Pegasus Mission Managerwith the detailed technical and operational tasksof the mission integration process are the PegasusMission Engineer, the system integration team,and the launch site team.

6.1.1 Orbital Mission Responsibilities

As the launch service provider, Orbital’sresponsibilities fall into five areas: 1) ProgramManagement, 2) Mission Management, 3) MissionEngineering, 4) Launch Site Operations, and 5)Safety.

6.1.1.1 Pegasus Program Management

The Pegasus Program Manager has directresponsibility for Orbital’s Pegasus Program. ThePegasus Program Manager is responsible for allfinancial, technical, and programmatic aspectsof the Pegasus Program. Supporting the PegasusProgram Manager are the Contract Manager,Pegasus Chief Engineer, and Launch ServicesDirector. All contractual considerations areadministered between the payload and PegasusContract Managers. The Pegasus Chief Engineeris responsible for all technical aspects of thePegasus launch vehicle, to include vehicleprocessing and launch operations. The Directorof Launch Services is responsible for managementof all activities associated with providing thePegasus launch service, to include the Pegasuslaunch manifest, customer interface and missionplanning. The Launch Service Director providesthe customer with the management focus toensure the specific launch service customer’sneeds are met. This individual assists theadministration of the contract by providing theContract Manager with technical evaluation andcoordination of the contractual requirements.

6.1.1.2 Pegasus Mission Management

The Pegasus Mission Manager is the Pegasusprogram single point of contact for all aspects of

a specific mission. This person has theresponsibility to ensure contractual commitmentsare met within schedule and budget constraints.The Pegasus Mission Manager will co-chair theMission Integration Working Groups (MIWGs)with the payload Mission Manager. The PegasusMission Manager’s responsibilities includedetailed mission planning, launch vehicleproduction coordination, payload integrationservices, mission-peculiar designs and analysiscoordination, payload interface definition, launchrange coordination, integrated scheduling, launchsite and flight operations coordination.

6.1.1.3 Pegasus Mission Engineering

The Pegasus Mission Engineer is responsible forall engineering and production decisions for aspecific mission. This person has overall technicalprogram authority and responsibility to ensurethat a vehicle is produced, delivered to theintegration site, and integrated to support a specificmission requirements. The Mission Engineersupports the Pegasus Mission Manager to ensurethat vehicle preparation is on schedule andsatisfies all payload requirements for launchvehicle performance.

6.1.1.4 Pegasus Mechanical Engineering

The Pegasus Mission Mechanical Engineer isresponsible for the mechanical interface betweenthe satellite and the launch vehicle. This personworks with the Pegasus Mission Engineer toverify mission specific envelopes are documentedand environments, as specified in the ICD, areaccurate and verified.

6.1.1.5 Pegasus Engineering Support

The Pegasus engineering support organization isresponsible for supporting mission integrationactivities for all Pegasus missions. Primary supporttasks include mission analysis, softwaredevelopment, mission-peculiar hardware designand testing, mission-peculiar analyses, vehicleintegration procedure development andimplementation, and flight operations support.

6.1.1.6 Pegasus Launch Site Operations

The Launch Site Manager is directly responsible



for launch site operations and facilitymaintenance. All work that is scheduled to beperformed at the Orbital launch site is directedand approved by the Pegasus Launch SiteManager. This includes preparation andexecution of work procedures, launch vehicleprocessing, and control of hazardous operations.All hazardous procedures are approved by theappropriate customer launch site safety manager,the launch range safety representative, the PegasusLaunch Site Manager, and the Pegasus SafetyManager prior to execution. In addition, PegasusSafety and Quality Assurance engineers arealways present to monitor critical and hazardousoperations. Scheduling of payload integrationwith the launch vehicle and all related activitiesare also coordinated with the Launch SiteManager.

6.1.1.7 Pegasus Systems Safety

Each of the Pegasus systems and processes aresupported by the Pegasus safety organization.Systems and personnel safety requirements arecoordinated and managed by the Safety Manager.The Safety Manager is primarily responsible forperforming hazard analyses and developingrelevant safety documentation for the Pegasussystem. The Safety Manager works closely withthe launch system development, testing, payloadintegration, payload and launch vehicleprocessing, and launch operations phases toensure adherence to applicable safetyrequirements. The Safety Manager interfacesdirectly with the appropriate government rangeand launch site personnel regarding launchvehicle and payload ground safety matters. TheSafety Manager assists the mission team withidentifying, implementing and documentingpayload and mission unique safety requirements.

6.2 Mission Integration Process

The Pegasus mission integration process ensuresthe launch vehicle and payload requirements areestablished and implemented to optimize bothparties needs. The Pegasus integration process isstructured to facilitate communication andcoordination between the launch vehicle andpayload customer. There are four major

components to the integration process; 1) thePegasus and payload mission teams, 2) TechnicalInterchange Meetings, 3) Mission IntegrationWorking Groups and 4) the readiness reviewprocess.

6.2.1 Mission Teams

The mission teams are established in the initialphase of the mission planning activity to create asynergistic and cohesive relationship betweenthe launch vehicle and payload groups. Theseteams consist of representatives from each of themajor disciplines from each group, i.e.,management, engineering, safety, and quality.The mission teams are the core of the integrationprocess. They provide the necessary continuitythroughout each phase of the integration processfrom initial mission planning through launchoperations. The team is responsible fordocumenting and ensuring the implementationof all mission requirements via the payload toPegasus Interface Control Document (ICD).

6.2.2 Integration Meetings

Two major types of meetings are used toaccommodate the free-flow of informationbetween the mission teams. The TechnicalInterchange Meeting (TIM) is traditionally reservedfor discussions focusing on a single technicalsubject or issue. While TIMs tend to focus ontechnical and engineering aspects of the missionthey may also deal with processing and operationsissues as well. They are typically held via teleconto accommodate multiple discussionopportunities and/or quick reaction. TIMdiscussions facilitate the mission team decisionprocess necessary to efficiently and effectivelyimplement mission requirements. They are alsoused to react to an anomalous or unpredictedevent. In either case, the results of the TIMdiscussions are presented in the MissionIntegration Working Group (MIWG) meetings.The MIWG provides a forum to facilitate thecommunication and coordination of missionrequirements and planning. MIWGs are usuallyheld in a meeting environment to accommodatediscussion and review of multiple subjects andface-to-face resolution of issues. Pre-established



agendas will be used to ensure all appropriatediscussion items are addressed at the MIWG.Launch Operations Working Groups (LOWG),Ground Operations Working Groups (GOWG),Range Working Groups (RWG) and SafetyWorking Groups (SWG) are all subsets of theMIWG process. Results of the MIWGs arepublished to provide historical reference as wellas track action items generated by the missionteams. The number and types of MIWGs variesbased on the mission unique requirements. Figure6-2 summarizes the typical working groupmeetings.

6.2.3 Readiness Reviews

Each mission integration effort contains a seriesof readiness reviews to provide the oversight andcoordination of mission participants andmanagement outside the regular contact of theMIWG environment. Each readiness review

ensures all organizations are in a position toproceed to the next major milestone. At aminimum, two readiness reviews are baselinedinto the integration process; 1) the MissionReadiness Review (MRR) and 2) the LaunchReadiness Review (LRR). The MRR is typicallyheld 1-2 weeks prior to shipping the spacecraft tothe integration facility. The MRR provides aprelaunch assessment of the launch vehicle,spacecraft, facilities, and range readiness forsupporting the integration and launch effort. TheLRR is typically conducted 1-3 days prior tolaunch. The LRR serves as the final assessment ofall organizations and systems readiness prior toconducting the launch operation. Due to thevariability in complexity of different payloadsand missions the content, quantity and scheduleof readiness reviews are tailored to support themission unique considerations.

6.3 Mission Planning and Development

Orbital will assist the customer with missionplanning and development associated withPegasus launch vehicle systems. These servicesinclude interface design and configurationcontrol, development of integration processes,launch and launch vehicle related analyses,facilities planning, launch campaign planning toinclude range services and special operations,and integrated schedules. Orbital will supportthe working group meetings described in thissection, and spacecraft design reviews.

6.3.1 Baseline Mission Cycle

The procurement, analysis, integration and testactivities associated with the Pegasus launch of apayload typically occur over a 24-30 monthbaseline mission cycle. This baseline schedule,detailed in Figure 6-3, is not meant to be a rigidstructure, but a template for effective missionmanagement and payload integration.Throughout this time, Orbital will work closelywith personnel from the customer and otherorganizations involved in the launch to ensure asuccessful mission. The schedule in Figure 6-3shows a typical 24 month mission. The baselinemission cycle includes:

• Mission management, document exchanges,

L-24 to L-8Months

L-18 to L-8Months

L-18 to L-6Months

L-6 to L-2Months

L-4 to L-1Months

MIWGs

RWGs

SWGs

GOWGs

LOWGs

Meeting PurposeTimeframe

Figure 6-2. Summary of Typical Working Groups.PEG044

• Establish Mission Requirements• Document Mission Requirements• Coordinate Test and Support

Requirements

• Establish Mission Range Requirements

• Document Mission Range Requirements

• Coordinate Range Test and Support Documentation

• Establish Mission Safety Requirements

• Document Mission Safety Requirements

• Coordinate Mission Safety Support Requirements

• Establish Mission Operations and Processing Requirements

• Document Mission Operations and Processing Requirements

• Coordinate Operations and Processing Support Requirements

• Establish Mission Launch Operations Requirements

• Document Mission Launch Operations Requirements

• Coordinate Launch Operations Support Requirements

Release 5.0

August 2000

6-5

Pegasus User's G

uide Section 6.0—

Mission Integration

L-MonthsActivity 28 27 26 25 24 23 22 21 20 19 18 17 16 15 14 13 12 11 10 9 8 7 6 5 4 3 2 1 L 1 2

PEG047

ATPMission Integration

Mission Analysis

Interface Development

Interface Control Document (ICD)

Payload Milestones (P/L Dependent)

Drawings

Integrated Procedures

Mass Properties

Range Documentation (UDS)

Launch Vehicle

Range

Flight Plan/Trajectory

Safety Process

Payload Safety Reviews

Safety Documentation

Operations Planning

Launch Checklist/Constraints

Meetings/Rehearsals

Program Reviews

Launch Vehicle Hardware Review

Readiness Review

Initial Launch Capability (ILC)

Mission Integration Working Groups

Operations Working Groups

FinalDraft

Mechanical and Electrical Coupled Loads

Post-FlightReport

Draft

PRD PRD MissionAnnex

OR

PSP OD

Preliminary Final

FInal

PDR

ILC

CDR PL Arrivalat VAB

Integrated Procedures

Final

Draft Final

Rehearsal

Mission Requirements

PreliminaryDrawings

Ground Safety Approval

Final Drawings

Final ARAR PreliminaryProcedures

Final Pro-edures

Draft ARAR

Kickoff Preliminary

Prelim Mass Props

Prelim Mass Props

Motor Pre-Ship Review MRR LRR

KEY:ATP - Authority to ProceedARAR - Accident Risk Assessment ReportCDR - Critical Design ReviewICD - Interface Control DocumentILC - Initial Launch CapabilityLRR - Launch Readiness ReviewMRR - Mission Readiness ReviewOD - Operations DirectiveOR - Operations Requirement DocumentPDR - Preliminary Design ReviewPRD - Program Requirements DocumentPSP - Program Support PlanUDS - Universal Document SystemVAB - Vehicle Assembly Building

Payload DocumentLaunch Vehicle DocumentMilestoneReview

Figure 6-3. T

ypical Mission C

ycle.



meetings and reviews required to coordinateand manage the launch service;

• Mission and payload integration analysis;

• Design, review, procurement, testing andintegration of all mission-peculiar hardware;and

• Range interface, safety, and launch site flightand operations activities and reviews.

6.4 Interface Design and Configuration Control

Orbital will develop a mission-unique payloadICD to define the interface requirements for thepayload. The ICD documents the detailedmechanical, electrical and environmentalinterfaces between the payload and Pegasus aswell as all payload integration specifics, includingground support equipment, interface testing andany unique payload requirements. The ICD isjointly approved by the customer and Orbital.An integrated schedule will also be developed.

6.5 Safety

Ground and flight safety is a top priority in anythe launch vehicle activity. Pegasus launchvehicle processing and launch operations areconducted under strict adherence to USgovernment safety standards. The lead range atthe integration and launch sites are the ultimateresponsibility for overall safety. These rangeshave established requirements to conduct launchvehicle and satellite processing and launchoperations in safe manner for both those involvedas well as the public. Launch vehicle and payloadproviders must work together with the rangesafety organizations to ensure all safetyrequirements are understood and implemented.

6.5.1 System Safety Requirements

In the initial phases of the mission integrationeffort, regulations and instructions that apply tospacecraft design and processing are reviewed.Not all safety regulations will apply to a particularmission integration activity. Tailoring the rangerequirements to the mission unique activities willbe the first step in establishing the safety plan.Pegasus has three distinctly different mission

approaches effecting the establishment of thesafety requirements:

1) Baseline mission: Payload integration andlaunch operations are conducted atVandenberg Air Force Base (VAFB), CA

2) Ferry mission: Payload integration isconducted at VAFB and launch operationsare conducted from a non-VAFB launchlocation.

3) Campaign mission: Payload integration andlaunch operations are conducted at a siteother than VAFB.

For the baseline and ferry missions, spacecraftprelaunch operations are conducted at Orbital’sVehicle Assembly Building (VAB), Building1555,VAFB. For campaign style missions, thespacecraft prelaunch operations are performedat the desired launch site.

Before a spacecraft arrives at the processing site,the payload organization must provide thecognizant range safety office with certificationthat the system has been designed and tested inaccordance with applicable safety requirements(e.g. EWR 127-1 Range Safety Requirements forbaseline and ferry missions). Spacecraft thatintegrate and/or launch at a site different than theprocessing site must also comply with the specificlaunch site’s safety requirements. Orbital willprovide the customer coordination and guidanceregarding applicable safety requirements.

Figure 6-4 provides a matrix of the governingsafety requirements for demonstrated and plannedPegasus payload integration flow. The Orbitaldocuments listed in the matrix closely follow theapplicable range safety regulations.

It cannot be overstressed that the applicablesafety requirements should be considered in theearliest stages of spacecraft design. Processingand launch site ranges discourage the use ofwaivers and variances. Furthermore, approval ofsuch waivers cannot be guaranteed.

6.5.2 System Safety Documentation

Range safety requires certification that spacecraft



systems are designed, tested, inspected, andoperated in accordance with the applicableregulations. This certification takes the form ofthe Missile System Pre-Launch Safety Package(MSPSP) (also referred to as the Accident RiskAssessment Report (ARAR)) which describes allhazardous systems on the spacecraft andassociated ground support equipment (GSE).Hazardous systems include ordnance systems,separation systems, solar array deploymentsystems, power sources, RF and ionizing radiationsources, and propulsion systems. The MSPSPmust describe all GSE used at the processing andlaunch sites, with special attention given to lifting,handling GSE, and pressurization or propellantloading equipment. EWR 127-1 Chapter 3Appendix 3A provides an outline of a typicalMSPSP.

At certain sites, specific approval must be obtainedfor all radiation sources (RF and ionizing). Orbitalwill coordinate with the spacecraft organizationand the specific site safety office to determinedata requirements and obtain approval. Datarequirements for RF systems normally includepower output, center frequency, scheduling timesfor radiating, and minimum safe distances. Datarequirements for ionizing sources normallyinclude identification of the source, source

VAFB

VAFB

CCAFB

KSC

VAFB

WFF

VAFB

VAFB

CCAFS

CCAFS

CCAFS

WFF

WFF

KMR

LaunchSite

Applicable SafetyRequirements Documents

Payload Integration Site

Figure 6-4. Applicable Safety Requirements.

PE

G04

5

EWR 127-1 / Orbital TD-0005 / Orbital TD-0018



EWR 127-1 / KHB 1710 / Orbital TD-0005 / Orbital TD-0018

EWR 127-1 / RSM-93 / Orbital TD-0005 / Orbital TD-0018

RSM-93 / Orbital TD-0005 / Orbital TD-0018

EWR 127-1 / KMR Range Safety Manual / Orbital TD-0005 / Orbital TD-0018

strength, half-life, hazard control measures, andminimum safe distances.

The MSPSP must also identify all hazardousmaterials that are used on the spacecraft, GSE, orduring operations at the processing and launchsites. Some examples of hazardous materials arepurge gases, propellant, battery electrolyte,cleaning solvents, epoxy, and adhesives. AMaterial Safety Data Sheet must be provided inthe MSPSP for each hazardous material. Also anestimate of the amount of each material used onthe spacecraft or GSE, or consumed duringprocessing shall be provided.

The MSPSP also shall specify the groundoperations flow and identify those operationsthat are considered hazardous. Hazardousoperations include lifting, pressurization, batteryactivation, propellant loading, and RF radiatingoperations.

All hazardous procedures that will be performedat the processing or launch site must be submittedto the specific site safety office for approval.Additionally, Orbital shall review and approvehazardous spacecraft procedures to ensurepersonnel at Orbital facilities will be adequatelyprotected from harm. Orbital shall provide thecoordination necessary for timely submission,review and approval of these procedures.

6.5.3 Safety Approval Process

Figure 6-5 depicts the typical safety approvalprocess for a commercial Pegasus mission. Ifpermitted by the processing and launch sitesafety organizations, it is recommended thattailoring of the applicable safety requirements beconducted early in the spacecraft design effort.This will result in greater understanding of thesite-specific regulations, and may provide moreflexibility in meeting the intent of individualrequirements. This is especially critical for newlydesigned hazardous systems, or new applicationsof existing hardware.

It is encouraged that safety data be submitted asearly as practical in the spacecraft developmentschedule. The review and approval processusually consists of several iterations of the MSPSP



and hazardous procedures to ensure allrequirements are met and all hazards areadequately controlled. Working sessions are heldperiodically to clarify the intent of requirementsand discuss approaches to hazard control. Theseworking sessions are normally scheduled tocoincide with existing Mission Integration

Figure 6-5. Safety Approval Process. PEG046

Identification ofApplicable

Requirements

Payload Organi-zation Incorporates

Orbital and SiteSafety Comments

Orbital SubmitsMSPSP and AnyComments to Re-quired Site Safe-ty Organizations

Working Sessionsto Tailor Specific

Requirements(if Required)

Working Sessionsas Required to

Review Comments

PayloadOrganization

Submits MSPSP toOrbital for Review

Have All CommentsBeen Adequately

Addressed?

No

Yes

MSPSPApproved

Working Groups and Ground Operation WorkingGroups.

When certain requirements cannot be satisfiedas specifically stated in the regulation, theapproving safety organization at the processingand launch sites may waive the requirementwhen provided sufficient justification. This requestfor variance must contain of an identification ofthe requirement, assessment of the risk associatedwith not meeting the letter of the requirement,and the design and procedural controls that arein place to mitigate this risk. As stated previously,the use of variances is discouraged and approvalcannot be guaranteed.

Section 7.0—Ground and Launch Operations


Pegasus User's Guide Section 7.0—Ground and Launch Operations

7.1 Pegasus/Payload Integration Overview

The Pegasus system has been designed tominimize both vehicle and payload handlingcomplexity as well as launch base operationstime. Horizontal integration of the Pegasusvehicle simplifies integration procedures,increases safety and provides excellent accessfor the integration team. In addition, simplemechanical and electrical interfaces and checkoutprocedures reduce vehicle and payloadintegration times, and increase system reliability.

7.2 Ground and Launch Operations

Figure 7-1 shows a typical ground and launchoperations flow which is conducted in threemajor phases:

• Launch Vehicle Integration: Assembly andtest of the Pegasus vehicle;

• Payload Processing: Receipt and checkoutof the satellite payload, followed byintegration with Pegasus and verification ofinterfaces; and

• Launch Operations: Mating of Pegasus withthe carrier aircraft, take-off and launch.

Each of these phases is more fully describedbelow. Orbital maintains launch site managementand test scheduling responsibilities throughoutthe entire launch operations cycle. Figure 7-2provides a typical schedule of the integrationprocess through launch.

7.2.1 Launch Vehicle Integration

7.2.1.1 Integration Sites

All major vehicle subassemblies are deliveredfrom the factory to the Vehicle Assembly Building(VAB) at Orbital’s integration sites. Orbital'sprimary integration site is located at VandenbergAir Force Base (VAFB), California. Through theuse of the OCA, this integration site can supportlaunches throughout the world. The VAFB OCAhotpad area is shown in Figure 7-3.

The following Pegasus GSE is maintained at theVAB:

Figure 7-1. Typical Processing Flow.PEG050

Prior to satellite arrival at the Vehicle Assembly Building (VAB) the Pegasus motors and avionics section are built up, integrated, and tested.

Upon arrival at the VAB the satellites are prepared for mating with the Pegasus.

Following satellite prepara-tions and interface checks the satellites are mated to Pegasus.

After completion of the satellite mate to Pegasus an integrated test is conducted to ensure compatibility between the satellites and Pegasus.

Once the satellites and Pegasus have successfully checked out the payload fairing is installed on Pegasus.

The integrated Pegasus and satellites are then transported to the Orbital Carrier Aircraft (OCA) and mated with the modified L-1011.

Final flightline preparations are performed with the Pegasus and satellites prior to launch.

The one hour captive carry portion of the launch operations provides the launch team with the final checkout of the Pegasus and satellites prior to launch.

Pegasus is launched from the OCA at an altitude of 39,000 feet and drops for 5 seconds before the first stage ignites.

Release 5.0

August 2000

7-2

Pegasus User's G

uide Section 7.0—

Ground and Launch O

perations

L-Weeks

L-Days7 6 5 4 3 2 1 0

Activity 15 14 13 12 11 10 9 8 7 6 5 4 3 2 1

PEG051

Motor Arrival at the VAB

Motor Build Up

Wing and Aft Skirt Installation

Avionics Section Arrival at the VAB

Avionics Section Testing

Flight Simulation 1

Motor Stages Mated

Flight Simulation 2

Payload Arrival at the VAB

Payload Preparations

Payload Interface Verification Test

Payload Electrical Mate to Pegasus

Flight Simulation 3

Payload Mechanical Mate to Pegasus

Flight Simulation 4

Pre-Fairing Closeout Activities

Faring Installation

Vertical Fin Removal

Transfer Pegasus to AIT

OCA Arrival at VAFB

Pegasus Mate to OCA

Combined Systems Test

Pre-Launch Prepreations

Launch

Figure 7-2. T

ypical Pegasus Integration and T

est Schedule.



• An Assembly and Integration Trailer (AIT),stationary rails, and motor dollies for serialprocessing of Pegasus missions.

• Equipment for transportation, delivery,loading and unloading of the Pegasus vehiclecomponents.

• Equipment for nominal integration and test ofa Pegasus vehicle.

• Equipment to maintain standard payloadenvironmental control requirements.

• General equipment to allow mating of thepayload with the Pegasus vehicle (Orbitaldoes not provide payload specific equipment).

7.2.1.2 Vehicle Integration and Test Activities

Figure 7-4 shows the Pegasus stages being

integrated horizontally at the VAB prior to thearrival of the payload. Integration is performed ata convenient working height, which allows easy

Figure 7-3. Orbital Carrier Aircraft Hot Pad Area at VAFB.

Non-HazardousOperations Area

HazardousOperations Area Taxiway

AsphaltShoulder

Stairs

61 m to Taxiway

Orbital Carrier AircraftGround PowerStation

GSETrailer

Lavatory

Wing JackWing Jack

Nose Jack

NitrogenTubeTruck

B-4 MaintenancePlatform

AIT

Exclusion Area/Stayout ZonePersonnel Access Restricted toPersons Specifically Identifiedin Work Package Procedures

Notes:

1 - Payload HEPA Filter2 - Air Conditioning Unit3 - Aircraft Ground Power Unit

Taxiway

AsphaltShoulder

CL

CL

Runway/Taxiway

AirfieldWindsock

0 15 30

Scale(Meters)

45 60

1

2 3

PEG054

Figure 7-4. Pegasus Integration.PEG055



access for component installation, test, andinspection. The integration and test processensures that all vehicle components andsubsystems are thoroughly tested before andafter final flight connections are made.

Vehicle systems tests include a series of tests thatverify operation of all subsystems prior to stagemate. The major tests are Vehicle Verification,Phasing Tests and Flight Simulations. For each ofthese a specialized test software load is installedinto the Pegasus Flight Computer.

Vehicle Verification is a test that efficientlycommands all subsystems (fin actuators, TVCs,FC discrete outputs, RCS, pyro commands, etc.)in an accelerated time line.

Phasing tests verify the sign of the control loop ofthe flight actuators and the dynamic operation ofthe IMU. In this test the IMU is moved manuallywhile the motion of the flight actuators (fins,TVCs and RCS) is observed and recorded.

Flight simulation testing uses the actual flightcode and simulates a “fly to orbit” scenario. Allflight actuators, pyro commands and FCcommands are exercised. The Flight Simulationis repeated after each major vehicle configurationchange (i.e., Flight Simulation #1 after motorstages are built-up, Flight Simulation #2 afterstage mate, Flight Simulation #3 after payloadelectrically mated/jumpered and Flight Simulation#4 after the payload is mechanically mated).After each test, the configuration of the vehicle isfrozen until a full and complete data review ofthe test is complete, which usually takes one totwo days. The payload nominally participates inFlight Simulation #3 and #4.

In addition to these major tests, several other testsare performed to verify the telemetry, flighttermination, accelerometer and RF systems.

Pegasus integration activities are controlled by acomprehensive set of Pegasus Work Packages(PWPs), which describe and document in detailevery aspect of integrating Pegasus and itspayload. Pegasus Mission Specific EngineeringWork Packages (EWPs) are created for missionunique or payload specific procedures.

7.2.2 Payload Processing

A typical Pegasus payload is delivered to theintegration site at launch minus 30 calendardays. The payload completes its own independentverification and checkout prior to beginningintegrated processing with Pegasus at the launchsite. Initial payload preparation and checkout isperformed by payload personnel prior to FlightSimulation #3.

Payload launch base processing procedures andpayload hazardous procedures should becoordinated through Orbital to the launch rangeno later than 120 days prior to first use (draft) and30 days prior to first use (final).

7.2.2.1 Ground Support Services

The payload processing area capabilities willdepend on which mission option is chosen basedon launch site – integrate and launch; integrate,ferry, and launch; or Pegasus campaign to launchsite.

Vandenberg ground support services whichwould be used in the launch and ferry scenariosare outlined in Appendix C.

7.2.2.2 Payload to Pegasus Integration

The integrated launch processing activities aredesigned to simplify final launch processing whileproviding a comprehensive verification of thepayload interface. The systems integration andtest sequence is engineered to ensure all interfacesare verified after final connections are made.

7.2.2.2.1 Pre-Mate Interface Testing

The electrical interface is verified using a missionunique Interface Verification Test (IVT), inconjunction with any payload desired testprocedures, to mutually verify that the interfacemeets specifications. The IVT and payloadprocedures include provisions for testing theLPO interfaces, if necessary.

If the payload provider has a payload simulator,this test can be repeated with this simulator priorto using the actual payload. These tests,customized for each mission, typically checkout



the LPO controls, launch vehicle sequencing,and any off-nominal modes of the payload.

When the payload arrives at the launch sitePegasus can be made available for a preliminarymechanical interface verification before finalpayload preparations.

After “safe-to mate” tests, the payload iselectrically jumpered, and further interface testing(e.g., data flow between the spacecraft and thePegasus) is performed, if necessary. FlightSimulation #3 is then performed, using a flightMDL, IMU simulator, and other EGSE. Forpayloads with simplified interfaces to the Pegasus,it may be acceptable to proceed to payload mateand the final Flight Simulation, immediately afterthe IVT.

7.2.2.2.2 Payload Mating and Verification

Once the payload aft end closeouts are completed,the payload will be both mechanically andelectrically mated to the Pegasus. Followingmate, the flight vehicle is ready for the finalintegrated systems test, Flight Simulation #4, inflight configuration. One of the last two flightsimulations is performed on the flight batteries.This test is in full flight configuration (internalpower, firing RCS, etc.), but without ordnanceconnected, allowing a complete check of allinterfaces after mating the payload, whileminimizing the payload time on the vehiclebefore launch. The integrated test proceduresare developed by the LOWG and reviewed bythe appropriate payload, launch vehicle andsafety personnel.

7.2.2.2.3 Final Processing and Fairing Close-Out

After successful completion of Flight Simulation#4, all consumables are topped-off and ordnanceis connected. Similar payload operations mayoccur at this time. Once consumables are topped-off, final vehicle/payload closeout is performedand the payload fairing is mated. Integratedsystem tests are conducted to ensure that thePegasus/payload system is ready for launch afterpayload mate.

7.2.2.2.4 Payload Propellant Loading

Payloads utilizing integral propulsion systemswith propellants such as hydrazine can be loadedand secured through coordinated Orbital,Government and payload contractorarrangements for use of the propellant loadingfacilities in the VAB. All launch integrationfacilities will be configured to handle these sealedsystems in the integration process with the launchvehicle. The propellant loading facility ismaintained visibly clean.

7.2.3 Launch Operations

7.2.3.1 Orbital Carrier Aircraft Mating

The Pegasus is transported on the Assembly andIntegration Trailer (AIT) to the OCA for mating.This activity typically takes place about threedays prior to launch. Once Pegasus is mated tothe OCA, Orbital monitors the Hot Pad 24 hoursper day through launch.

The OCA/LPO/Pegasus interface is fully verifiedprior to mating the launch vehicle to the carrieraircraft by performing an OCA Pre-Mate ElectricalCheckout. Mission unique/payload LPO Stationinterfaces are also verified using a mission specificEWP prior to Pegasus mate to the OCA. Using theAIT, the Pegasus ground crew then mates thevehicle to the OCA.

All OCA/LPO/Pegasus/payload interfaces are thenverified again through a functional test, know asthe combined systems Test (CST). The CST alsoverifies the interfaces with the range tracking,telemetry, video and communications resources.If the payload has an arming plug which inhibitsa pyrotechnic event, and this plug was not installedin the VAB, it may be installed at this timethrough the fairing access door.

The payload can continue to maintain access tothe payload through this door up to one hourprior to aircraft engine start (approximately take-off minus two hours). After engine #2 start, theground air conditioning system is removed andthe fairing environment is thermally controlledby the AACS from the aircraft, which flows intothe fairing under the control of the LPO.



7.2.3.2 Pre-Flight Activities

The pre-departure activities and launch checklistflow is shown in Figure 7-5. The first procedurefor the mission operations team begins after therange communications checks and setup at take-off (T.O.) minus 4.5 hrs. At T.O. minus 3.5 hrs,the LPO enters the carrier aircraft and powers upPegasus upon direction from the LaunchConductor (LC). Concurrently, final closeout ofPegasus is accomplished and the range safetyengineers verify that the FTS is functioning bysending arm and fire commands to the FTSantennas via actual range assets or a range testvan.

Other Pegasus verification tests are thenperformed to exercise most aspects of the Pegasus,ensuring the vehicle will switch from carrieraircraft power to internal battery power and thatthe IMU, flight computer, and telemetry systemare all working correctly. Payload operations are

verified to ensure the payload can be controlledby the LPO control switches as required. End-to-end checks are made to verify Pegasus andpayload (if applicable) telemetry transmissionsare received in the telemetry room.

7.2.3.3 Launch Control Organization

The Launch Control Organization normallyconsists of three separate groups. TheManagement group includes the MissionDirectors for the launch vehicle and the payloadand a senior Range representative. The OrbitalMission Director provides the final PegasusProgram recommendation for launch decisionbased on inputs from the Vehicle Engineer andthe Launch Conductor. Similarly, the PayloadMission Director polls the various payloadpersonnel to determine the readiness of thepayload for launch, and the Range respresentativeprovides the final Go/No-Go for the Range.

Figure 7-5. Typical Pegasus Launch Checklist Flow.

TO - 4.5 Hrs.

TO - 4.25 Hrs.

TO - 4 Hrs.

TO - 3.5 Hrs.

TO - 2 Hrs.

TO- 1.5 Hrs.

TO - 30 Min.

TO - 5 Min.

L-58 Min

L-30 Min

L-4 Min

50 MinuteBuilt-In Hold

15 MinuteBuilt-In Hold

FTS Open Loop Test

Power System Test

Pegasus Power-Up

Range/Facility Setup Climb/Cruise ImmediateHazard

EmergencyJettison

Internal Power

Terminal Count

Post LaunchOCA RTB

ControlledJettisonRTB

Recycle

Pre-Takeoff Poll

OCA Taxi

Engine Start

PEG053

Pre-Engine Start

Ground Operations Captive Carry Flight Contingency

Abort



The second group is the Operations/EngineeringGroup, including the Launch Conductor, theVehicle Engineer and the Range Control Officers.The Orbital Launch Conductor is responsible forrunning the countdown procedure. The OrbitalVehicle Engineer has the overall responsibilityfor the Pegasus launch vehicle. A team ofengineers, which reviews the telemetry to verifythe system is ready for launch, support the VehicleEngineer. The range status is coordinated by theRange Control Officer who provides a Go/No-Go status to the Launch Conductor.

The third group is the Airborne Operations Groupwhich includes the Launch Panel Operator (LPO)and the aircraft crew. The LPO monitors on-board systems from the launch panel stationonboard the carrier aircraft and executes on-board countdown procedures. The aircraft crewoperates the aircraft, achieves proper pre-releaseflight conditions and activates the actual physicalrelease of the Pegasus vehicle.

7.2.3.4 Flight Activities

The launch checklist begins prior to engine startand continues until after Pegasus is released. Allmembers of the launch team and the aircraftcrew work from this procedure. Abort proceduresand emergency procedures are also contained inthe launch notebooks.

At the hotpad about one hour before take-off, theFTS power is turned on and all inhibits areverified, the S&A safing pins are removed, andthe vehicle is placed in a ready state. At this timethe aircraft and the Pegasus are ready for take-off.

Orbital arranges for Pegasus telemetry andtracking services during captive carry and Pegasuspowered flight. Data will be passed to thepayload mission control console as determinedby the LOWG and MIWG process.

Once airborne, Pegasus is configured into alaunch condition by switching the FTS to internalbattery power at L-15 min, the avionics bus tointernal power at L-6 min, and the transientpower bus to internal power at L-4 min. If theLPO station is supplying external power to the

spacecraft, the spacecraft will be transitioned tointernal power no later than L-6 minutes. At L-45sec, the fin thermal batteries are activated and asinusoidal fin sweep is commanded by the flightcomputer to all fins to verify that they are workingcorrectly. The fin sweep telemetry, fin positionand command current, are monitored and, if theyare nominal, the Pegasus is “Go For Launch.”The Orbital Launch Conductor relays this "Go"from the Pegasus control center to the pilotcommander. After confirmation from the pilotcommander of a go for launch, the LaunchConductor performs the drop countdown. Thepilot releases Pegasus on the Launch Conductor'scommand. After release, the Pegasus flight isautonomous with the exception of the positivecommand capability for flight termination in theevent of an anomalous flight.

7.2.3.5 Abort/Recycle/Return-to-BaseOperations

The approximate time to recycle in the air is 30minutes. The minimum stand-down time after anabort/return-to-base is 24 hours. Orbital plansand schedules all required contingency landingareas and support services prior to each launchattempt. In general, only minimal support servicesare available to the payload at contingencylanding sites. Available recycle time is dependenton payload constraints as well. For example, thepayload must determine battery margins to verifyrecycle capabilities. Payload providers mustspecify the maximum time they can withstandthe absence of GSE support.

Section 8.0—Documentation


Pegasus User's Guide Section 8.0—Documentation

8.1 Interface Products and Schedules

Orbital divides external interfaces into two areas:interfaces with the Pegasus production team (i.e.,our subcontractors and vendors), typically forhardware products, and interfaces with externalorganizations, which are typically documentationproducts and data exchanges.

External organizations with which Orbital willhave information exchanges include the launchvehicle customer, the payload provider, the range,and the U.S. Department of Transportation. Theproducts associated with these organizations areincluded within the 24 to 30-month baselinePegasus mission cycle. As such, Orbital referencesrequired dates in a “launch minus” timeframe.The major products and submittal times associatedwith these organizations are divided into twoareas — those products that Orbital produces,detailed in Figure 8-1, and those products thatare required by Orbital, detailed in Figure 8-2.

8.2 Mission Planning Documentation

The available Pegasus documentation includes acollection of formal and informal documentsdeveloped and produced by Orbital. The numberof separate formal documents required for asuccessful mission has been minimized byconsolidation of documents and maximizing theinformal exchange of information (e.g., workinggroups) before inclusion on formal, controlledconfiguration documents such as the payloadInterface Control Document (ICD).

8.3 Mission-Unique Analyses

Mission analysis, which includes trajectory/GN&C analyses and environment analyses,begins shortly after mission authorization isreceived. Orbital will generate the optimaltrajectory to the desired orbit, determine theguidance parameters, and evaluate the autopilotstability. From these analyses, the Mission DataLoad (MDL) will be generated and then tested inreal-time simulations.

8.3.1 Trajectory Analysis

Orbital will perform a Preliminary and FinalMission Analysis using POST and the Orbital-

developed Non-Real Time Simulation (NRTSim)analysis tool, which performs six degree-of-freedom simulations. The primary objective is todetermine the compatibility of the payload with

Preliminary ICDPreliminary Mission Analysis/Mission ProfilePreliminary Trajectory AnalysisFinal ICD

Final Mission Analysis/Mission ProfilePost-Flight ReportProgram RequirementsDocumentPRD Mission AnnexPegasus Flight TerminationSystem ReportPegasus Accident RiskAssessment ReportPreliminary MissionConstraints DocumentPreliminary Launch ChecklistOperations RequirementsDocumentPreliminary TrajectoryFinal TrajectoryFinal Launch ChecklistMission Constraints DocumentLaunch Specific Flight PlanPayload DescriptionVehicle Information Message

ProductDelivered(L-Weeks)

Deliveredto

Customer

Range

Department ofTransportation

Figure 8-1. Documentation Produced by Orbital forCommercial Pegasus Launch Services.

ATP+90 Days*L-74

L-52Customer Sign+30 Days**L-12

L+7L-52

L-52As Required

As Required

L-26

L-26L-22 (Ops-60 Days)L-9L-6L-6L-6L-9L-9L-2

PE

G05

6

* Or Prior to Payload PDR ** Or Prior to Payload CDR

Mission Unique Services DefinitionMission Requirements SummaryPreliminary Payload Drawing/Mass PropertiesPayload PRD InputFinal Payload DrawingPayload Accident Risk AssessmentReportChecklist/Launch Constraint InputsIntegration ProceduresFinal Payload Mass PropertiesProgram Support PlanOperations DirectiveFlight Plan Approval

ProductDueDate

(L-Weeks)

Deliveredby

Figure 8-2. Documentation Required by Orbital forCommercial Pegasus Launch Services.

ATPL-92L-76

L-56L-48L-45

L-28L-24L-8L-48L-15L-1

Customer

Range

PE

G05

7



Pegasus and to provide succinct, detailed missionrequirements, such as payload environments,performance capability, accuracy estimates andpreliminary mission sequencing. Much of thedata derived from the Preliminary MissionAnalysis is used to establish the ICD and performinitial range coordination.

Orbital will perform recontact analysis for post-separation events to determine if a C/CAM isrequired. The analysis will verify that sufficientseparation distance exists between the payloadand final stage following payload separation andwill include effects of separation system operationand residual final stage thrust.

8.3.2 Guidance, Navigation and ControlAnalyses

Consists of several separate detailed analyses tothoroughly evaluate the planned mission and itseffects throughout powered flight. The trajectorydesign, guidance, stability, and control analysesresult in a verified mission-unique flight softwareMDL.

Guidance Analysis — Pegasus dispersions andinjection accuracies are determined usingpredicted vehicle motor performance, massuncertainties, and aerodynamic and INS errors.Uncertainties are combined to obtain estimateddispersions in perigee, apogee, inclination andargument of perigee. This data is incorporated inthe payload ICD.

Stability and Control Analysis — Using theoptimum trajectory from POST, Orbital selects aset of points throughout Stage 1 burn forinvestigating the stability characteristics of theautopilot. For the exo-atmospheric portions offlight, the autopilot margins are similarly evaluatedat discrete points to account for the changingmass properties of the vehicle. The controlsystem gains are chosen to provide adequatestability margins at each operating point. Orbitalvalidates these gains through perturbed flightsimulations designed to stress the functionality ofthe autopilot and excite any possible instabilities.Due to the proprietary nature of Orbital’s controlalgorithms, this analysis is not a deliverable to the

payload vendor.

8.3.3 Coupled Loads Analysis

Orbital has developed finite element structuralmodels of the Pegasus vehicle. Orbital willperform a coupled loads analysis to determinemaximum responses of the entire stack. A singleload cycle is run after a payload modal surveyhas taken place and a test verified payload modelhas been supplied. The coupled loads analysiswill also contain a "rattlespace analysis." Thisanalysis verifies the payload does not violate thepayload fairing dynamic envelope.

8.3.4 Payload Separation Analysis

Orbital will use the Pegasus STEP simulation toensure that the payload is in the desired orientationfor successful separation at the end of boost.Orbital will perform a separation tip-off analysisto verify the three axis accelerations that thepayload will experience during the separationevent from the final stage. This analysis will onlybe conducted on an Orbital-supplied separationsystem.

8.3.5 RF Link and Compatibility Analyses

RF link analyses will be updated for each trajectoryto ensure sufficient RF link margins exist for boththe telemetry and flight termination systems.

8.3.6 Mass Properties Analysis and Mass DataMaintenance

Orbital will track and maintain all mass properties,including inertias, relating to the Pegasus vehicle.Payload-specific mass properties provided toOrbital by the customer will be included. Allflight components are weighed prior to flight andactual weights are employed in final GN&Canalyses. Orbital will require estimates of thepayload mass to facilitate preliminary missionplanning and analyses. Final payload massproperties are required at least 75 days prior tolaunch within the tolerances specified.

8.3.7 Power System Analysis

Orbital develops and maintains a power budgetfor each mission. A mission power budget will



verify that sufficient energy and peak load marginexist. Battery usage is strictly controlled on thevehicle and batteries are charged prior to vehicleclose-out.

8.3.8 Fairing Analyses

Two payload-specific analyses performed byOrbital relate to the payload fairing. These are acritical clearance analysis (contained in thecoupled loads analysis) based on the dimensionsand payload characteristics provided by thecustomer and a separation point analysis toselect the timing for this event. Payload fairingmaximum deflection occurs at pull-up.

The fairing separation point is nominally timed tocoincide with dynamic pressure falling below0.01 psf which usually occurs during the Stage 2burn. Payload requirements specifying lowerdynamic pressures or aerodynamic heatingenvironments at fairing deployment may beaccommodated by delaying this separation event.In general, this separation delay will lead to somedegradation in Pegasus payload performance,which will need to be evaluated on a case by casebasis.

8.3.9 Mission-Unique Software

Mission-unique flight software consists of theflight MDL, which contains parameters andsequencing necessary to guide Pegasus throughthe desired trajectory.

Prior to each flight, Orbital evaluates theinteraction of the flight MDL with the mission-independent guidance and control software inthe Guidance and Control Lab (GCL). Orbitalpersonnel conduct a formalized series ofperturbed trajectories, representing extremedisturbances, to ensure that both the flight MDLand the G&C software are functioning properly.MDL performance is judged by the ability of thesimulation to satisfy final stage burnoutrequirements. The final flight MDL verification isobtained by conducting a closed-loop real-timesimulation.

8.3.10 Post-Launch Analysis

Orbital will provide a detailed mission report to

the customer normally within six weeks of launch.Included in the mission report will be the actualtrajectory, event times, environments and otherpertinent data as reduced from telemetry fromonboard sensors and range tracking. Orbital alsoanalyzes telemetry data from each launch tovalidate Pegasus's performance.

8.4 Interface Design and Configuration Control

Orbital will develop a mission-unique payloadICD to succinctly define the interfacerequirements for the payload. This documentwill detail mechanical, electrical andenvironmental interfaces between the payloadand Pegasus as well as all payload integrationspecifics, including ground support equipment,interface testing and any unique payloadrequirements. The customer and Orbital jointlyapprove the ICD.

Section 9.0—Shared Launch Accommodations


Pegasus User's Guide Section 9.0—Shared Launch Accommodations

Orbital has extensive experience in integratingand launching multiple payloads. Multiplespacecraft configurations have been flown onover half of the Pegasus missions to date.

Two technical approaches are available foraccommodating multiple payloads. These designapproaches are:

Load-Bearing Spacecraft — aft spacecraftdesigned to provide the structural load pathbetween the forward payload and the launchvehicle, maximizing utilization of available massperformance and payload fairing volume

Non Load-Bearing Spacecraft — aft spacecraftwhose design cannot provide the necessarystructural load path for the forward payload

9.1 Load-Bearing Spacecraft

Providing a load-bearing aft payload maximizesuse of available volume and mass. The availablemass for the aft payload is determined by thePegasus performance capability to orbit less theforward payload and attach hardware mass. Allremaining mission performance, excluding a stackmargin, is available to the aft payload. The load-bearing spacecraft interfaces directly to Pegasusand the forward payload via pre-determinedinterfaces. These interfaces include standardOrbital separation systems and pass-throughelectrical connectors to service the forwardpayload. Figure 9-1 illustrates this approach.

Two approaches may be taken for load-bearingspacecraft. The first approach involves the use ofan Orbital design using the MicroStar bus,successfully developed and flown for ORBCOMMspacecraft. The MicroStar bus features a circulardesign with an innovative, low-shock separationsystem. The spacecraft bus is designed to allowstacking of co-manifested payloads in "slices"within the fairing. The bus design is compact andprovides exceptional lateral stiffness.

The second approach is to use a design developedby other spacecraft suppliers, which must satisfyPegasus and forward payload structural designcriteria. The principal requirements levied uponload-bearing spacecraft are those involving

mechanical and electrical compatibility with theforward payload. Structural loads from the forwardpayload during all flight events must be transmittedthrough the aft payload to the Pegasus. Orbitalwill provide minimum structural interface designcriteria for shear, bending moment, axial andlateral loads, and stiffness.

For preliminary design purposes, coupled effectswith the forward payload can be considered as arigid body design case with Orbital providedmass and center of gravity parameters. Integratedcoupled loads analyses will be performed withtest verified math models provided by the payloadcontractors. These analyses are required to verifythe fundamental frequency and deflections of thestack for compliance with the Pegasus requirementof 20 Hz minimum. Design criteria provided byOrbital will include “stack” margins to minimizeinteractive effects associated with potential designchanges of each payload. Orbital will providethe necessary engineering coordination betweenthe spacecraft and launch vehicle.

Electrical pass-through harnesses will also needto be provided by the aft payload along with

Figure 9-1. Load-Bearing Spacecraft Configuration.

TypicalForward

SpacecraftVolume

TypicalAft Load Bearing

SpacecraftVolume

(2.95)

φ 46.00(Dynamic)

(3.95)

FairingDynamicEnvelope

φ 23.00Sep Ring

φ 38° Avionics Thrust Tube(22.00" Long)

φ 38°Separation Ring

PEG058


Pegasus User's Guide Section 9.0—Shared Launch Accommodations

provisions for connectors and interfaceverification. The spacecraft supplier will need toprovide details of the appropriate analyses andtest to Orbital to verify adequacy of margins andshow that there is no impact to the forwardspacecraft or the launch vehicle.

9.2 Non Load-Bearing Spacecraft

For aft spacecraft that are not designed forwithstanding and transmitting structural loadsfrom the forward payload, the flight-proven DualPayload Attach Fitting (DPAF) is available on anoptional basis.

The DPAF structure (Figure 9-2) is an all graphitestructure which provides independent load pathsfor each satellite. The worst-case “design payload”for the DPAF is a 193 kg (425 lbs) spacecraft with51 cm (20 in) center of mass offset and first lateralfrequency of 20 Hz. The DPAF is designed toaccommodate this “design payload” at both theforward and aft locations, although the combinedmass of the two payloads cannot exceed Pegasuscapabilities. The upper spacecraft loads are

Primary Payload Volume

+Y

+X

φ 76.0 29.9

φ 66.026.0

Ogive Radius 269.2 106.0

101.640.0

102.940.5

Primary Payload Separation Plane

φ 114.345.0

DynamicEnvelope

Pegasus Avionics

Adapter Core Separation PlaneBeginning of Ogive

Secondary Payload Separation Plane

43 Separation17 System

97 Separation 38 System

128.850.7

Figure 9.2. Dual Payload Attach Fitting Configuration.

55.922.0

AvailableSecondary

PayloadVolume

5823

Separation System

Dimensions in cmin

PEG059

transmitted around the lower spacecraft via theDPAF structure, thus avoiding any structuralinterface between the two payloads.

The DPAF uses an Orbital standard 58 cm (23 in)Marmon clamp band interface for the upperpayload mounted on a separable adapter conewhich provides the transition to the 97 cm (38 in)cylinder. The aft satellite support structure consistsof a 43 cm (17 in) separation system and a 43 cm(17 in) adapter cone which transitions to the 97cm (38 in) diameter Pegasus third stage.

The separation systems are aluminum Marmonclamp designs. Each satellite is provided anindependent electrical interface to the launchvehicle including zero-force connectors tominimize tip-off at deployment.

The separation sequence for the stack beginswith initiation of the forward payload separationsystem followed by the separation of the conicaladapter. The aft payload is then separated andejected from within the cylinder which remainswith the third stage.

Section 10.0—Non-Standard Services

Release 5.0 July 2000 10-1

Pegasus User's Guide Section 10.0—Non-Standard Services

This section describes optional non-standardservices available. The earlier non-standardservice requirements are identified, the better,preferably at Orbital ATP. Many of these non-standard services have flight heritage on one ormore Pegasus flights.

10.1 Additional Fairing Access Doors

Additional access doors are available. Standardsizes are 8.5” x 13” and 4.5” circular. Certainrestrictions apply to door location.

10.2 Alternative Integration Sites

As a non-standard service, Pegasus can use thefollowing sites for payload integration:

• Eastern Range;

• Wallops Flight Facility; and

• Other sites are possible and will beinvestigated on a case-by-case basis and mayrequire inter-governmental coordination.

Pegasus will be integrated at Vandenberg andflown to the alternate integration site. The Pegasuswill be demated from the OCA, transported to theintegration facility, the fairing will be removed,payload integration activities will be conducted,the fairing will be reinstalled, and the Pegasuswill be transported back to the OCA and preparedfor launch.

10.3 Alternative Range Services

As a non-standard service, specifically to supporttrajectories not attainable without significanttrajectory dog-leg from Vandenberg, the Pegasuscan be launched from the following ranges:

• Eastern Range;

• Wallops Flight Facility; and

• Other ranges are possible and will beinvestigated on a case-by-case basis andmay require inter-governmental coordina-tion.

This assumes that the rocket and payloadintegration takes place at Vandenberg and the

integrated launch vehicle/satellite is ferried to thelaunch site on the OCA and launched withoutdemating from the OCA.

10.4 Class 10,000 Fairing Environment

Orbital can provide payload fairing purge withair meeting FED-STD-209E Class 10,000 (M5.5),in accordance with TD-0289, “PegasusContamination Control Plan, NASA Class M5.5Missions.” This task includes installing, operating,monitoring, and cleaning special HEPA-andcarbon-filtered conditioned-air supply systemsduring four phases of integrated operations:

• Inside the integration facility (VehicleAssembly Building);

• During transport to Hot Pad;

• During Hot Pad ground operations; and

• During Orbital Carrier Aircraft matedoperations.

10.5 Class 10,000 Payload/Vehicle IntegrationEnvironment

Orbital can provide a payload/vehicle integrationenvironment that is clean, certified, andmaintained at FED-STD-209E Class 10,000 (M5.5)to support payload mate through fairing closeoutoperations. This includes assembly of the payloadseparation system in a Class 10,000 cleanroom,final assembly of the payload fairings and othercomponents in Class 10,000 cleanrooms,preparation and verification of the Class 10,000softwall cleanroom, and integration of the Pegasusand payload in the cleanroom. All these tasksinclude cleaning operations, monitoring,verifying, and recording cleanliness levelsfrequently.

10.6 Fairing Internal Surface Cleaning

Orbital can clean, certify, and maintain internalsurfaces of the Pegasus payload fairing to MIL-STD-1246C, Level 750A, 600A or 500A. Thisinvolves increased levels of precision cleaning ofthe internal fairing surfaces prior to payloadencapsulation; additional surface cleanlinessmeasurements to verify surface cleanliness; and



additional handling controls to maintaincleanliness.

10.7 40-Pin Pass-Through Harness

As a non-standard service, Pegasus canincorporate 20 twisted shielded pairs of wiresfrom the payload interface plane to the OCA.This wiring matches the specifications of thestandard 5 pass-through pairs: 22 gauge wire,90% shielding, 2.5 ohms resistance, and amaximum carrying capability of 3.0A per wirepair.

10.8 Hydrazine Auxiliary Propulsion System

The Hydrazine Auxiliary Propulsion System(HAPS) improves injection accuracy (Table 3.1)and increases payload capability above 600 kmby 25 to 120 kg. Contact Orbital to obtain theexact performance capability associated with theHAPS. HAPS is more effective at higher altitudeand also permits injection of shared payloadsinto different orbits. HAPS is available as anoptional enhancement to Pegasus.

HAPS, which is mounted inside the AvionicsStructure, consists of a hydrazine propulsionsubsystem and a Stage 3 separation subsystem.After burnout and separation from the Stage 3motor, the HAPS hydrazine thrusters provideadditional velocity and both improvedperformance and precise orbit injection.

The HAPS propulsion subsystem (Figure 10-1)consists of a centrally mounted tank containingapproximately 59 kg (130 lbm) of hydrazine,helium pressurization gas, and three fixed, axiallypointed thrusters. The hydrazine tank containsan integral bladder which will support multiplerestarts.

10.9 Hydrocarbon Monitoring

Orbital can provide continuous monitoring ofhydrocarbon levels during all integrated payload/Pegasus operations. This requires the installation,calibration and frequent round-the-clockmonitoring of fixed and portable hydrocarbon(VOC) detectors in the Vehicle Assembly Building,during rollout to Hot Pad, and during Hot Pad

operations through fairing closeout. Also requiredare computer-controlled contamination datarecorders and alarming systems, for continuouscapture of hydrocarbon level data and remotewarning of excessive levels.

10.10 Instrument Purge System

Orbital can provide an instrument purge systemcapable of delivering GN2 or GHe to a payload.Orbital’s quick disconnect system exerts lessthan 50 lbf on the payload fitting. Orbital alsooffers the ability to cycle the system up to aflowrate of 40 SCFM with a pressure dropdownstream of the Pegasus/OCA interface of lessthan 75 psi. After fairing closeout on the Hot Padthe payload is limited to a maximum of 703 kg(1550 lbm) of GN2 or 50 kg (110 lbm) of GHethroughout captive carry.

10.11 Load Isolation System

Orbital can provide a Load Isolation System thatwill lower the fundamental frequencies of thepayload to avoid dynamic coupling with thePegasus fundamental frequencies at drop. ThisLoad Isolation System will decrease volume andmass available to the payload, to be quantified bythe frequency modification requirements of thepayload.

10.12 Low Tip-Off Rate Payload Attach Fittings

Clamp band separation impulse is one of theprimary causes of tip-off on the Pegasus separationsystem. Reduced Marmon clamp tension ispossible for some payloads that are significantlybelow the structural capabilities of the separationsystem.

10.13 Downrange Telemetry Support

Orbital has established relationships with anumber of government organizations to providetelemetry coverage beyond the capability of thelaunch-range fixed telemetry assets. These mobileassets can be deployed in advance to anappropriate down range location or in near real-time (airborne systems) to support the acquisitionof telemetry from either Pegasus or spacecraft(spacecraft telemetry downlink dependent)



Figure 10-1. Hydrazine Auxiliary Propulsion System (HAPS).PEG060

Rocket EngineAssemblies (3)

+Z

+X

Hydrazine TankFill/Drain Valve (3)

Pyro Isolation

Valve

70.7427.85

Side View

Isometric View

φ 100.9739.75

Dimensions incmin



telemetry. These systems have been usedsuccessfully on a number of Pegasus missionsand prove to be a cost-effective means ofcollecting telemetry for real-time re-transmissionor for post-flight data review. Orbital willcoordinate spacecraft requirements with themobile range provider to ensure appropriateoperational support and data products areprovided to the payload customer.

10.14 Payload Connector Covers

Flight-proven connector covers are available forthe payload side of the separation system tocover the 42-pin and 18-pin interface connectors.The connector covers are spring loaded andattach to the standard umbilical support brackets.At payload separation, the spring-loadedaluminum cover snaps closed covering theconnector.

10.15 Payload Fit Check Support

Pegasus can send flight and non-flight hardwareand test support personnel to the payloadcontractor site for a fit check. Support hardware(flight fairing, flight or universal frangible jointdepending on when fit check is to be performedand payload contractor’s ability to supportordnance operations, mock Stage 2/3 interstage,inert Stage 3, mock avionics section) and technicaland engineering support will be sent to thepayload contractor’s designated site to support afairing fit check with flight hardware.

10.16 Payload Propellant Loading

Orbital can provide for full hydrazine or bi-propellant loading services. This service can beperformed in the Pegasus Vehicle AssemblyBuilding at Vandenberg AFB, CA.

10.17 Pegasus Separation System Test Unit

Orbital can provide a Pegasus Separation SystemTest Unit (PSSTU) and Avionics Structure to thepayload contractor. The PSSTU is a non-flightseparation system that is provided to payloadcontractors to perform pyroshock characterizationtesting. The pyroshock test plan should besubmitted to Orbital 30 days prior to testing for

Orbital concurrence on the use of the PSSTU andAvionics Structure. The PSSTU and AvionicsStructure will be delivered to the spacecraftcontractor two weeks prior to the required needdate for pyroshock testing and returned to Orbitalno later than two working days after the conclusionof pyroshock testing. Orbital will review andcheck the test set up prior to firing the bolt cuttersfor pyroshock testing. Orbital must witness thetest.

The PSSTU may not be used by the payloadcontractor to perform any testing other thanpyroshock characterization and may not be usedas a spacecraft build/transportation fixture.Electrical harnessing and connectors for thePSSTU are the responsibility of the payloadcontractor and will not be supplied by Orbital.Contractor must identify need date of PSSTU atleast six months prior to need date.

10.18 Round-the-Clock Payload Support

Pegasus supports a nominal eight-hour per day,five day per week work schedule prior to payloadfairing mate. During certain launch vehicleoperations, hours will be briefly exceeded. Facilitysafety requirements dictate that Orbital employeesmust be present during payload processing. As anon-standard service, payload supportrequirements prior to payload fairing mate outsidethese hours can be satisfied.

10.19 Serial Telemetry Interface

Orbital offers a polled payload Serial TelemetryInterface to incorporate payload telemetry datainto the Pegasus launch vehicle telemetrydownlink. The RS-422/485 interface employs aserial link between the payload and the Pegasusflight computer. The flight computer interrogatesthe payload at a predetermined rate and receivespayload data to be interleaved into thedownlinked telemetry stream. The serial streamis radiated to the ground via the Pegasus S-Bandnetwork. The telemetry data volume cannotexceed 125 bytes/sec.

10.20 Spin Stabilization Above 60 RPM

As a non-standard service, Orbital can provide



the necessary supplies and services to separatepayloads into a spin stabilization mode above 60rpm, the nominal limit.

10.21 Stage 2 Onboard Camera

Pegasus can fly a real-time second stage videosystem. This self-contained system has adedicated battery, RF signal transmission system,and two cameras for forward and aft views of therocket. The cameras switch views as commandedby the flight computer to capture critical stagingevents and fairing separation. It can also beswitched from the LPO control station while incaptive carry.

10.22 State Vector Transmission From Pegasus

As a non-standard service, Pegasus can utilizethe serial telemetry link with the payload totransmit a state vector from the flight computerdirectly to the satellite. This state vector will bein a format specified in the Pegasus SerialTelemetry Specification. State vector accuracywill be that of the Pegasus inertial navigationsystem.

10.23 Thermal Coated Forward SeparationRing

Prior to separation system assembly, Orbital canprovide the customer a forward payloadseparation system ring for application of thermalcoating or thermal blankets. All work proceduresand added materials must be approved by Orbitalin advance of ring shipment.

Appendices

Release 5.0 August 2000 A-1

Pegasus User's Guide Appendix A—Payload Questionnaire

A Payload Questionnaire (PQ) is required fromthe payload organization for use in preliminarymission analysis. The PQ is the initialdocumentation of the mission cycle and is needed

at least 22 months before the desired launch date.It is not necessary to fill out this PQ in its entiretyto begin mission analysis. Simply provide anyavailable information.

Spacecraft Name

Spacecraft Owner

Spacecraft Manufacturer

Spacecraft Purpose

Point of Contact

Nominal Launch Date

Launch Window Description

Mission Timeline

Final Orbit Apogee

Final Orbit Perigee

Final Orbit Inclination

Maximum Apogee Allowable

Minimum Perigee Allowable

Argument of Perigee

Right Ascension of Ascending Node

Apogee Accuracy

Perigee Accuracy

Inclination Accuracy

Argument of Perigee Accuracy

Right Ascension of Ascending Node Accuracy

km

km

deg

km

km

deg

deg

km

km

deg

deg

deg

nmi

nmi

deg

nmi

nmi

deg

deg

nmi

nmi

deg

deg

deg

Spacecraft Information

SI UnitsParameter

Trajectory Requirements

Mission Information

English Units



Propellant Type, Orbit Insertion

Propellant Type, Station Keeping

Multiple Burn Capability?

Propellant Mass

Specific Impulse

Spacecraft Mass (Maximum or Nominal)

Spacecraft Coordinate System

C.M. - Thrust Axis (Origin at Interface Plane)

C.M. - Y Axis

C.M. - X Axis

C.M. - Z Axis

C.M. Tolerance - Thrust Axis

C.M. Tolerance - Y Axis

C.M. Tolerance - Z Axis

Coefficients of Inertial - Ixx

Coefficients of Inertial - Ixx Tolerance

Coefficients of Inertial - Iyy

Coefficients of Inertial - Iyy Tolerance

Coefficients of Inertial - Izz

Coefficients of Inertial - Izz Tolerance

Coefficients of Inertial - Ixy

Coefficients of Inertial - Ixy Tolerance

Coefficients of Inertial - Iyz

Coefficients of Inertial - Iyz Tolerance

Coefficients of Inertial - Ixz

Coefficients of Inertial - Ixz Tolerance

Y/N

lbm

sec

lbm

in

in

in

in

in

in

slug ft2

slug ft2

slug ft2

slug ft2

slug ft2

slug ft2

slug ft2

slug ft2

slug ft2

slug ft2

slug ft2

slug ft2

slug ft2

±

±

±

±

±

±

±

±

±

±

±

±

±

±

±

±

±

±

SI UnitsParameter

Propulsion

English Units

Mass Properties

Y/N

kg

sec

kg

mm

mm

mm

mm

mm

mm

mm

kg m2

kg m2

kg m2

kg m2

kg m2

kg m2

kg m2

kg m2

kg m2

kg m2

kg m2

kg m2

The Coefficients of Inertia Below Are About the Center of Mass



Spacecraft Height

Spacecraft Diameter

Dimensional Drawing/CAD Model

Payload Separation System Supplier

Payload Adapter Supplier

Fairing Access Door Locations

Mission Specific Hardware

Pegasus Provided Pryo Commands

Pegasus Provided Discrete Commands

Payload EMI/RFI Susceptibility

Source

Function

Role

Band

Frequency

Bandwidth

Power Output

Sensitivity

Prelaunch Temperature Range

Prelaunch Relative Humidity Range

Maximum Prelaunch Gas Impingement Velocity

Maximum Ascent Heat Flux

Maximum Free-Molecular Heat Flux

Maximum Fairing Ascent Depressurization Rate

mm

mm

dbµ V/m

C

°C

%

m/sec

W/m2

W/m2

mbar/sec

in

in

MHz

°F

%

ft/sec

BTU/hr ft2

BTU/hr ft2

psi/sec

SI UnitsParameter

Mechanical Interface

English Units

Electrical Interface

Payload RF Emitters and Receivers

Electromagnetic Compatibility

Thermal Environment



Maximum Allowable Flight Accuracies

Allowable Acoustics Curve

Maximum Allowable Sine Vibration

Allowable Sine Vibration Curve

Maximum Allowable Shock

Allowable Shock Curve

Maximum Lateral Acceleration

Maximum Longitudinal Acceleration

Fundamental Frequency - Lateral

Fundamental Frequency - Longitudinal

Fairing Air Cleanliness

Maximum Deposition on Spacecraft Surfaces

Outgassing - Total Weight Loss

Outgassing - Volatile Condensable Material

Weight Loss

Maximum Allowable Tip-Off Rate

Spin-Up Required at Separation?

Desired Spin-Up Rate

Pointing Requirement

Maximum Allowable Pointing Error

dB OA

Grms

g

g

g

Hz

Hz

Class

mg/m2

%

%

deg/sec

Y/N

rpm

deg

dB OA

Grms

g

g

g

Hz

Hz

Class

grains/ft2

%

%

deg/sec

Y/N

rpm

deg

SI UnitsParameter

Dynamic Environment

English Units

± ±

Contamination Control

Orbital Injection Conditions



The following questions pertain to Pegasus Launch Operations and should be provided to Orbital assoon as possible after contract start:

Flightline Operations

1. Provide a brief description of any testing to be performed at the flightline on the day of launchoperations:

2. What is the maximum expected duration of the testing?

❑ <30 minutes

❑ <60 minutes

❑ >60 minutes (provide further detail)

3. Will the testing involve GSE or ASE?

❑ GSE

❑ ASE

4. Provide a brief description of types of closeouts expected at the flightline on the day of launchoperations

❑ Mechanical:

❑ Electrical:

❑ Software:

5. What is the total maximum expected duration of these closeouts?

❑ <30 minutes

❑ <60 minutes

❑ >60 minutes (provide further detail)

6. Specify any transition of spacecraft control/monitor functions from GSE or ASE?

7. Provide a brief description of any timers or restrictions associated with flightline closeouts (e.g.,battery plugs, solar array deployment, etc.)



8. Specify payload LPO readback actions required during captive carry:

❑ Telemetry:

❑ Power Supply:

❑ Heaters:

❑ Other (specify):

9. Is telemetry available to ground or LPO or both?

❑ LPO

❑ Ground

10. Describe any final configuration functions the payload LPO must perform during captive carry(e.g., keyboard input commands, power down payload trickle charge, etc.):

Safety Operations

11. Are there any unique LPO safety monitor systems?

❑ Yes (provide description)

❑ No

Power Down/Power Up

12. Provide a brief description of Spacecraft configuration steps in the event Pegasus cycles powerduring ground operations:

Abort Operations

13. In the event of an abort, describe any payload LPO re-configuration operations (e.g., batterytrickle charge power up, etc.):



14. In the event of an abort, is there any GSE required immediately upon landing?

15. In the event of a return to remote landing site, are there any unique GSE transportation issues?

Release 5.0 August 2000 B-1

Pegasus User's Guide Appendix B—Electrical Interface Connectors

1.0 Wiring

Orbital provides one 42-pin umbilical harnessdedicated for payload use. The standard interfaceconnects the payload to the Pegasus flightcomputer as well as to the Launch Panel OperatorStation located in the carrier aircraft. All wiringshall be 22 AWG. Twisted Shielded Pair (TSP)passthroughs shall not exceed 3 A current perwire pair.

The standard connector is configured as shownin Figure B-1.

2.0 Connectors

Figure B-2 defines the pin assignments for thestandard payload interface connector at theseparation plane. The connectors are as follows:

Launch vehicle side: 42 pin plug with pin contacts:

MS-27484T-16F-42P

Payload side: 42 pin receptacle with socket

contacts:

MS-27474T-16F-42S

Orbital will provide the payload contractor withthe payload half of the electrical separationconnectors for integration into the payloadharness.

3.0 Non-Standard Interfaces

Depending on the mission, non-standardinterfaces may still be accommodated on theinterface connectors by taking advantage ofunused functions.

Figure B-1. Standard Payload Electrical Connec-tions.

5 Payload Passthrough Pairs

1 RS-422 Bi-Directional Serial Interface

4 Discrete Talkback Inputs (Breakwire-Type) to Pegasus Flight Computer

8 Discrete Commands from Pegasus Flight Computer to Payload

1 Payload Separation Sense to Pegasus Flight Computer

1 Spare Wire Pair

Connector Function Allocation

10

4

8

16

2

2

Number of Wires

PE

G06

1

Figure B-2. Payload Interface Connector Pin Assignments for P-65/J-2 Connector.

Pin Name Function Standard Destination

PEG062A

1 PPT1 + Payload Passthrough 1 + LPO Station

2 PPT1 - Payload Passthrough 1 -









11 CMD1 + Discrete Command 1 + FC Discrete Output 9

12 CMD1 - Discrete Command 1 -








Release 5.0 August 2000 B-2

Pegasus User's Guide Appendix B—Electrical Interface Connectors

Figure B-2. Payload Interface Connector Pin Assignments for P-65/J-2 Connector (continued).

Pin Name Function Standard Destination

PEG062B








27 P/L SEP + Payload Separation Sense + FC Discrete Input 10

28 P/L SEP - Payload Separation Sense -

29 TB1 + Discrete Talkback 1 + FC Discrete Input 5

30 TB1 - Discrete Talkback 1-


32 TB2 - Discrete Talkback 2 -





37 TLM TXD + RS-422/485 TXD + FC Serial Channel 12

38 TLM TXD - RS-422/485 TXD -

39 TLM RXD + RS-422/485 RXD +

40 TLM RXD - RS-422/485 RXD -

41 Spare Spare N/A

42 Spare Spare

Release 5.0 August 2000 C-1

Appendix C—VAFB VehiclePegasus User's Guide Assembly Building Capabilities

1.0 Ground Support Services

The payload processing area within the VAB willbe made available to the payload 30 calendardays prior to launch for independent payloadcheck-out. This area is intended to allow payloadpreparations prior to mate.

All work performed within the VAB is scheduledthrough the Orbital Site Manager. Orbital willsupport and schedule all payload hazardous orRF test operations conducted within the VABwhich require Range notification or approval.

2.0 Payload Servicing Areas

The VAB includes a payload preparation areaaccessible via motorized roll-up doors and doubledoors. Personnel access is via separate doors.Separate areas in the facility are designated forpayload servicing, test, and integration withsufficient space for payload-specific checkoutequipment.

The VAB is temperature and humidity controlledand kept "visibly clean." A soft-walled cleanroom is available if required for cleanliness levelsgreater than visibly clean for payload preparationand mating. The cleanroom will enclose PegasusStage 3 during processing as shown in Figure C-1. Floor loading is consistent with a fully loadedPegasus on its AIT.

3.0 Available Ground Support Equipment

The VAB is equipped with 552 Kpa (80 psi)compressed air and 115 VAC/220 VAC 3 phasepower. Overhead sodium lamps provide aminimum of 824 lux (75 ft-candles) of illumination

in the payload and vehicle processing areas. Fulllightning protection and dedicated extendedbuilding grounding comply with the standardsfor ordnance processing. Conductive floor surfaceand continuous grounding strips support the fullbuilding and personnel antistatic disciplines.

All personnel are required to wear leg stats whenworking near the rocket in the high bay areas ofthe VAB. Access to the integration facility isstrictly controlled with a badging system. Thenumber of payload personnel allowed in theentire facility is limited to no more than 10 at anytime whenever Pegasus motors are in the facility.This requirement will vary depending on totalfacility activities and is driven by operationalsafety constraints.

Orbital will provide a forklift, hydraulic lift table,5-ton bridge crane, and 1-ton cleanroom cranefor payload handling, as needed. Any payloadspecific handling hardware required forinterfacing with the lift table or crane (e.g. handlingcrane, rotation fixture, attachments, testequipment, etc.) should be supplied by thepayload unless other arrangements have beenmade.

4.0 Payload Work Areas

Orbital will provide approximately 37 m2 (400ft2) of work space in the west coast VAB forpayload use starting 30 calendar days prior to aplanned launch operation and extending to oneweek after launch. Approximately 9 m2 (100 ft2)of administrative office space will be provided ata site close to the VAB.

Release 5.0 August 2000 C-2

Appendix C—VAFB VehiclePegasus User's Guide Assembly Building Capabilities

Figure C-1. The Vandenberg Vehicle Assembly Building General Layout.

AirComp

GSE Storage Area

Wom

en’s

Bat

hroo

m

Zone WestHVAC/

Utility Room

Bat

tery

/E

lect

rical

Lab

Securityand

AccessControl

Site Manager’sPlanning

Area

ConfRoom

ElectricalTechs

MechanicalTechs

Engineers andTechnical Staff

OperationsPlanning Area

West BayVehicle Processing

557m2

6,000 ft2

East BayVehicle Processing

557 m2

6,000 ft2

36118

15.250

15.250

36118

HazardousPropellant Loading

Area

MGSEReceipt andInspection

Ground ElectronicsSupport Area

NASAProgram Office

Work AreaMen’s

Bre

akR

oom

Payload

Work Area"B"

Flight ComponentBonded Storage

Area

*CypherBlast Door

*CypherBlast Door

4.314

Sliding Door

W

US

AF

Pro

gram

Offi

ce W

ork

Are

a

* = Cypher Door

2.48 W x H3.0

104.615 W x

Roll-Up Door

Roll-Up Door

Roll-Up Door

Roll-Up Door

H3.712

7.625 W x H6.1

20

7.625 W x H6.1

20

4.6

15W

xH

3.7

12

Roll-Up Door

6.120 W x H6.1

20

Roll-Up Door

6.120 W x H6.1

20

*CypherBlast Door

Men’sBathroom

Haz OpControlRoom

Em

er.

Die

sel

Gen

.

Hal

lway

Ele

ctric

al S

yste

ms

and

Coo

ling

Zon

esW

est

Eas

t

4.615

9.832

16.554

Blast Door

Blast Door

Soft Wall Clean Room

Soft Wall Clean Room

16.5

54

0.250.83

ConcreteBlast Wall

Window

Double WideDoor

Window

Wom

en’s

*

*

PayloadWork Area

"A"

** * *

*

Com

m. S

ysT

elep

hone

High PressureGas SystemControl Area

Planning QABreakRoom

Pas

sage

way

Pay

load

Wor

kA

rea

"C"

4.916 x

PayloadArea

6.120

Dimensions in mft

PEG064

Release 5.0 August 2000 D-1

Pegasus User's Guide Appendix D—Launch Range Information

1.0 Introduction

Pegasus's air-launched design vastly increaseslaunch point flexibility. Some ground support isrequired to insure the safety of the people andproperty, to communicate with the carrier aircraftand to provide data collection and display. Thissupport is usually provided by a federal MajorRange and Test Facility Base (MRTFB) such as theEastern Range, Patrick AFB, FL; Western Range,Vandenberg AFB, CA; and Wallops Flight Facility,VA.

Pegasus has also been supported by the WallopsMobile Range for launch from foreign soil suchas from the Canary Islands, Spain. The use of acertified mobile range satisfies requirements ofthe Department of Transportation to enable alicensed commercial launch.

To assist customers who may wish to launch froma specific geographic location, this Appendix Dsummarizes the capabilities needed. This supportcould be provided by any facility meeting thefollowing requirements:

2.0 Range Safety

2.1 Trajectory Analysis

The planned trajectory must be analyzed todetermine if any populated areas will be overflownand if the risk is acceptable. Impact limit linesmust be developed to insure that the instantaneousimpact point (IIP) of any stage or debris does notimpact inhabited land. Reference the Easternand Western Range, Range Safety RequirementsDocument (EWR 127-1) for detailed requirementsand risk limitations.

2.2 Area Clearance and Control

The airspace surrounding the launch area mustbe cleared and controlled during the mission.Notices to airmen and mariners must be sent toclear the airspace and the predicted impact pointsof the spent stages and known debris.

2.3 Range Safety Displays

Visual display of the present position and IIPsmust be available to the safety personnel to verifythat no safety criteria are violated. This requires

redundant tracking sources such as radar ortelemetry guidance data. Pegasus is equippedwith a C-Band tracking transponder and providesposition data in the telemetry downlink.

2.4 Flight Termination System

Pegasus is equipped with command receiversthat operate at either 416.5 or 425.0 MHz. Theyare capable of receiving commands utilizing thestandard four tone alphabet. The commandtransmitter system must meet federal standards asdescribed in EWR 127-1.

2.5 FTS Controllers

Certified FTS Controllers must meet the federalstandards described in EWR 127-1.

3.0 Telemetry

Pegasus downlinks telemetry data in the S-bandand upper S-band frequency range (2,200-2,300and 2,300-2,400 Mhz). A telemetry system mustbe capable of tracking, receiving and recordingthis data. The OCA has onboard video camerasand this data is transmitted via a telemetry systemthat operates in the upper S-band range. A chaseaircraft is normally used and it also downlinkstelemetry. A separate telemetry system is requiredto track, receive and record this data.

4.0 Communications

4.1 Air to Ground

Air to ground communications are required tocommunicate with the carrier aircraft during thelaunch operations. This can be in the HF, VHF orUHF frequency range.

4.2 Voice Nets

Voice nets are required for communicationsbetween the various controllers involved in theoperation. Four to eight nets are required.

5.0 Control Center

The launch team requires a control center toconduct the launch countdown. This centerrequires a minimum of twenty consoles withvoice nets and video displays. The consoles musthave the capability to remote key the radios for

Release 5.0 August 2000 D-2

Pegasus User's Guide Appendix D—Launch Range Information

communications with the carrier and chaseaircraft.

6.0 Data Requirements

6.1 Realtime Data

Realtime telemetry data must be provided toOrbital computers for logging and conversion tovideo displays. This data is used to monitor thehealth and status of the rocket and payload.

6.2 Video Distribution System

A video distribution system is required capable ofdisplaying a minimum of twelve video screens.

6.3 Recording

Recording of all the telemetry downlinks isrequired.

6.4 IRIG Timing

IRIG timing is required.

6.5 Weather Forecasts

Weather forecasts are required.

7.0 Optional Launch Ranges

Figure D-1 summarizes the additional launchranges available for Pegasus use, along with theinclinations that are achievable from each range.In addition, Orbital can as an optional service,launch Pegasus XL to low inclination easterlyorbits from alternative launch sites.

Figure D-1. Optional Launch Ranges and Achiev-able Inclinations.

Western Range (Baseline)

Eastern Range (Option)

Wallops Flight Facility (Option)

Alcantara (Future)

Kwajalein (Future)

Mission Unique Location (Requires Mobile Range)

(1)A broader range of inclinations may be achievable from each point, subject to additional analyses and coordination with range authorities. Additionally, lower inclinations than those indicated for each range can be achieved through dog-leg trajectories, with a commensurate reduction in performance. Some specific inclinations within these ranges may be limited by stage impact point or other restrictions.

Note:

70° to 130°

28° to 51°

38° to 55°

Equatorial

Equatorial

To Be Determined

RangeAchievable

Inclinations(1)

(Direct)

Established LaunchSites

Alternative LaunchSites

PEG065

Release 5.0 July 2000 E-1

Pegasus User's Guide Appendix E—Pegasus Flight History

FlightNumber Customer(s) Payload Payload Mission

Target OrbitLaunch Date Vehicle Type

320.0 x 360.0 nm @ 94.00° i

389.0 x 389.0 nm @ 82.00° i

405.0 x 405.0 nm @ 25.00° i

400.0 x 400.0 nm @ 70.00° i

450.0 x 450.0 nm @ 82.00° i

400.0 x 400.0 nm @ 70.00° i

195.0 x >1000 nm @ 70.02° i

398.0 x 404.0 nm @ 70.00° i

195.0 x >1000 nm @ 70.02° i

450.0 x 443.0 nm @ 90.00° i

298.0 x 394.0 km @ 97.13° i

340.0 x 955.0 km @ 97.40° i

350.0 x 4200.0 km @ 83.00° i

510.0 x 550.0 km @ 38.00° i

587.0 x 587.0 km @ 151.01° i

310.0 x 400.0 km @ 98.21° i

800.0 x 800.0 km @ 70.00° i

430.0 x 510.0 km @ 45.00° i

825.0 x 825.0 km @ 45.00° i

580.0 x 580.0 km @ 97.75° i

600.0 x 650.0 km @ 97.88° i

• Flight Test Instrumentation• Atmospheric Research• Communications Experiment• Tactical Communications Network

• Data Communications• Communications Experiment

• Technology Validation• Communications Experiment

• Technology Validation



• Communications• Atmospheric Research




• Atmospheric Research

• Space Physics Research



• Ocean Color Imaging



• LEO Communications

• University Science Payload• Commercial Telecommunica- tions Test Payload• Space Physics Research

PegaSat

SECS7 MicroSats

SCD-1OXP-1

ALEXISOXP-2

STEP-2

STEP-1

APEX(PegaStar)

FM1 &FM2MicroLab

STEP-3

REX-2

MSTI-3

TOMS

FAST

SAC-BHETE

MINISAT 01

OrbView-2

FORTE

STEP-4

8 ORBCOMM Satellites

SNOEBATSAT (T-1)

TRACE

4/5/90

7/19/91

2/9/93

4/25/93

5/19/94

6/27/94

8/3/94

4/3/95

6/22/95 3/8/96

5/16/96

7/2/96

8/21/96

11/4/96

4/21/97

8/1/97

8/29/97

10/22/97

12/23/97

2/25/98

4/1/98

Standard

Standard w/HAPS

Standard

Standard

Standard w/HAPS

XL

Standard

Standard (Hybrid)

XL XL

Standard (Hybrid)

XL

XL

XL

XL

XL

XL

XL

XL w/HAPS

XL

XL

DoD/NASA

DoDDoD

INPE BrazilOrbital

DoD/DoEOrbital

DoD

DoD

DoD

ORBCOMMNASA

DoD

DoD

BMDO

NASA

NASA

NASA

INTA Spain

Orbital/NASA

DoD

DoD

ORBCOMM-1

NASATeledesic

DoD/NASA

XF1

XF2

F3

F4

F5

F6

F7

F8

F9

F10

F11

F12

F13

F14

F15

F16

F17

F18

F19

F20

F21

Release 5.0 July 2000 E-2

Pegasus User's Guide Appendix E—Pegasus Flight History

FlightNumber Customer(s) Payload Payload Mission

Target OrbitLaunch Date Vehicle Type

818.5 x 818.5 km @ 45.02° i

818.5 x 818.5 km @ 45.02° i

750.0 x 750.0 km @ 25.00° i

635.0 x 700.0 km @ 70.00° i

540.0 x 540.0 km @ 97.56° i

550.0 x 550.0 km @ 97.75° i 775.0 x 775.0 km @ 97.75° i 825.0 x 825.0 km @ 45.02° i

405.0 x 1,750.0 km @ 69.00° i



• Data Communications• Atmospheric Experiment



• University Science Payload• Technology Validation


• Military Technology Demonstration



SCD-2Wing Glove

SWAS

WIRE

TERRIERSMUBLCOM


TSX-5

XL w/HAPS

XL w/HAPS

Standard (Hybrid)

XL

XL

XL w/HAPS

XL w/HAPS

XL w/HAPS

8/2/98

9/23/98

10/22/98

12/5/98

3/4/99

5/17/99

12/4/99

6/7/00

ORBCOMM-2

ORBCOMM-3

INPE BrazilNASA

NASA

NASA

NASAOrbital SSG

ORBCOMM-4

Orbital SSG/DoD

F22

F23

F24

F25

F26

F27

F28

F29

Pegasus

Documents

pegasus users guideapproved

pegasus users guide1

pegasus xl vehicle description

orbital sciencescorporation

orbital carrier aircraft

thepegasus launch system

payload interface systems

general performance