Orbital Pegasus User Manual
April 2010
Release 7.0
Pegasus® User’s Guide
Approved for Public Release
Distribution Unlimited
Copyright© 1996-2010 by Orbital Sciences Corporation. All Rights Reserved.
This Pegasus® User’s Guide is intended to familiarize potential space launch vehicle users with the
Pegasus launch system, its capabilities and associated services. The launch services described herein
are available for commercial procurement directly from Orbital Sciences Corporation.
Readers desiring further information on Pegasus should contact us via:
E-mail to: baldwin.bryan@orbital.com
Telephone: (703) 433-6043
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
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INTRODUCTION .................................................................................................................................... 1
1.
2. PEGASUS DESCRIPTION..................................................................................................................... 3
2.1. Pegasus XL Vehicle Description .................................................................................................... 3
2.1.1. Solid Rocket Motors .................................................................................................................. 3
2.1.2. Payload Fairing.......................................................................................................................... 3
2.1.3. Avionics .....................................................................................................................................4
2.1.4. Flight Termination System......................................................................................................... 5
2.1.5. Attitude Control Systems ........................................................................................................... 5
2.1.6. Telemetry Subsystem................................................................................................................ 6
2.1.7. Major Structural Subsystems.....................................................................................................6
2.1.7.1. Wing .................................................................................................................................... 6
2.1.7.2. Aft Skirt Assembly ............................................................................................................... 7
2.1.7.3. Payload Interface Systems ................................................................................................. 7
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TABLE OF CONTENTS
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2.2. Orbital Carrier Aircraft..................................................................................................................... 7
3. GENERAL PERFORMANCE CAPABILITY ........................................................................................... 8
3.1. Mission Profiles ..............................................................................................................................8
3.2. Performance Capability .................................................................................................................. 9
3.3. Trajectory Design Optimization ...................................................................................................... 9
3.4. Orbit Insertion Accuracy ................................................................................................................. 9
3.4.1. Actual Pegasus Insertion Accuracies ...................................................................................... 10
3.4.2. Error-Minimizing Guidance Strategies..................................................................................... 10
3.5. Collision/Contamination Avoidance Maneuver............................................................................. 12
4. PAYLOAD ENVIRONMENTS............................................................................................................... 13
4.1. Design Loads................................................................................................................................ 13
4.2. Payload Testing and Analysis ...................................................................................................... 14
4.3. Payload Acceleration Environment .............................................................................................. 14
4.4. Payload Random Vibration Environment .....................................................................................14
4.5. Sinusoidal Vibration...................................................................................................................... 15
4.6. Payload Shock Environment ........................................................................................................ 15
4.7. Payload Acoustic Environment..................................................................................................... 15
4.8. Pressure Profile ............................................................................................................................ 16
4.9. Payload Thermal Environment .....................................................................................................16
4.9.1. Payload Processing................................................................................................................. 16
4.9.2. Transportation ......................................................................................................................... 16
4.9.3. Ground Operations at the Flightline and Launch Operations.................................................. 21
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Powered Flight......................................................................................................................... 21
4.9.4.
4.9.5. Nitrogen Purge ........................................................................................................................ 21
4.10. Payload Electromagnetic Environment ........................................................................................ 21
4.11. Payload Contamination Control.................................................................................................... 22
5. SPACECRAFT INTERFACES.............................................................................................................. 23
5.1. Payload Fairing............................................................................................................................. 23
5.1.1. Fairing Separation Sequence .................................................................................................. 24
5.1.2. Payload Design Envelopes...................................................................................................... 24
5.1.3. Payload Access Door .............................................................................................................. 24
5.2. Payload Mechanical Interface and Separation System................................................................ 24
5.2.1. Standard Nonseparating Mechanical Interface ....................................................................... 24
5.2.2. Standard Separating Mechanical Interface ............................................................................. 24
5.3. Payload Electrical Interfaces ........................................................................................................ 29
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5.3.1. Standard Electrical Interface ...................................................................................................33
5.3.2. Mission-Unique Electrical Interface ......................................................................................... 34
5.3.2.1. Range Safety Interfaces/Vehicle Flight Termination......................................................... 35
5.3.2.2. Electrical Isolation ............................................................................................................. 35
5.3.2.3. Pre-Drop Electrical Safing................................................................................................. 35
5.4. Payload Design Constraints ......................................................................................................... 36
5.4.1. Payload Center of Mass Constraints....................................................................................... 36
5.4.2. Final Mass Properties Accuracy.............................................................................................. 36
5.4.3. Payload EMI/EMC Constraints................................................................................................ 36
5.4.4. Payload Stiffness..................................................................................................................... 37
5.4.5. Payload Propellant Slosh ........................................................................................................37
5.4.6. Customer Separation System Shock Constraints ................................................................... 37
5.4.7. System Safety Constraints ...................................................................................................... 38
5.5. Carrier Aircraft Interfaces ............................................................................................................. 38
5.5.1. Payload ASE Provisions..........................................................................................................38
5.5.2. Payload Support at Launch Panel Operator Station ...............................................................39
6. MISSION INTEGRATION..................................................................................................................... 39
6.1. Mission Management Structure....................................................................................................39
6.1.1. Orbital Mission Responsibilities............................................................................................... 39
6.1.1.1. Pegasus Program Management ....................................................................................... 40
6.1.1.2. Pegasus Mission Management ......................................................................................... 40
6.1.1.3. Pegasus Mission Engineering........................................................................................... 40
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TABLE OF CONTENTS (CONTINUED)
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6.1.1.4.
6.1.1.5. Pegasus Systems Safety .................................................................................................. 41
6.2. Mission Integration Process ......................................................................................................... 41
6.2.1. Mission Teams......................................................................................................................... 41
6.2.2. Integration Meetings ................................................................................................................ 42
6.2.3. Readiness Reviews ................................................................................................................. 42
6.3. Mission Planning and Development............................................................................................. 43
6.3.1. Baseline Mission Cycle............................................................................................................ 43
6.4. Interface Design and Configuration Control ................................................................................. 43
6.5. Safety .......................................................................................................................................... 43
6.5.1. System Safety Requirements .................................................................................................. 43
6.5.2. System Safety Documentation ................................................................................................ 45
6.5.3. Safety Approval Process .........................................................................................................45
7. Pegasus/Payload Integration Overview ............................................................................................... 46
7.1. Ground and Launch Operations ................................................................................................... 46
7.1.1. Launch Vehicle Integration......................................................................................................47
7.1.1.1. Integration Sites ................................................................................................................ 47
Pegasus Launch Site Operations ..................................................................................... 41
7.1.1.2. Vehicle Integration and Test Activities .............................................................................. 49
7.1.2. Payload Processing................................................................................................................. 49
7.1.2.1. Ground Support Services.................................................................................................. 50
7.1.2.2. Payload to Pegasus Integration ........................................................................................ 50
7.1.2.2.1. Pre-Mate Interface Testing......................................................................................... 50
7.1.2.2.2. Payload Mating and Verification ................................................................................. 50
7.1.2.2.3. Final Processing and Fairing Close-Out..................................................................... 50
7.1.2.2.4. Payload Propellant Loading ....................................................................................... 50
7.1.3. Launch Operations .................................................................................................................. 51
7.1.3.1. Orbital Carrier Aircraft Mating ........................................................................................... 51
7.1.3.2. Pre-Flight Activities ...........................................................................................................51
7.1.3.3. Launch Control Organization ............................................................................................ 51
7.1.3.4. Flight Activities .................................................................................................................. 52
7.1.3.5. Abort/Recycle/Return-to-Base Operations ....................................................................... 53
8. DOCUMENTATION.............................................................................................................................. 53
8.1. Interface Products and Schedules ...............................................................................................53
8.2. Mission Planning Documentation .................................................................................................54
8.3. Mission-Unique Analyses ............................................................................................................. 54
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Trajectory Analysis .................................................................................................................. 54
8.3.1.
8.3.2. Guidance, Navigation, and Control Analyses.......................................................................... 54
8.3.3. Coupled Loads Analysis ..........................................................................................................55
8.3.4. Payload Separation Analysis................................................................................................... 55
8.3.5. RF Link and Compatibility Analyses........................................................................................ 55
8.3.6. Mass Properties Analysis and Mass Data Maintenance ......................................................... 55
8.3.7. Power System Analysis ...........................................................................................................55
8.3.8. Fairing Analyses ...................................................................................................................... 55
8.3.9. Mission-Unique Software.........................................................................................................55
8.3.10. Post-Launch Analysis .............................................................................................................. 56
8.4. Interface Design and Configuration Control ................................................................................. 56
8.5. Mission Planning Schedule .......................................................................................................... 56
8.6. Payload Documentation Support.................................................................................................. 56
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9. SHARED LAUNCH ACCOMMODATIONS .......................................................................................... 56
9.1. Load-Bearing Spacecraft.............................................................................................................. 56
9.2. Non-Load-Bearing Spacecraft...................................................................................................... 57
10. NONSTANDARD SERVICES............................................................................................................... 58
10.1. Alternative Integration Sites ......................................................................................................... 58
10.2. Alternative Launch Sites............................................................................................................... 59
10.3. Downrange Telemetry Support .................................................................................................... 59
10.4. Additional Fairing Access Doors .................................................................................................. 59
10.5. Optional Payload/Vehicle Integration Environment...................................................................... 60
10.6. Enhanced Fairing Environment....................................................................................................60
10.7. Enhanced Fairing Internal Surface Cleaning ...............................................................................60
10.8. Hydrocarbon Monitoring ............................................................................................................... 60
10.9. Instrument Purge System............................................................................................................. 60
10.10. Increased Capacity Payload-to-GSE Interface ............................................................................ 61
10.11. Improved Insertion Accuracy Options .......................................................................................... 61
10.12. Load Isolation System .................................................................................................................. 63
10.13. Low Tip-Off Rate with Reduced Clamp Band Tension................................................................. 63
10.14. Enhanced Telemetry Capabilities – Payload Data ....................................................................... 63
10.15. State Vector Transmission From Pegasus................................................................................... 63
10.16. Payload Electrical Connector Covers........................................................................................... 64
10.17. Payload Fit Check Support........................................................................................................... 64
10.18. Payload Propellant Loading ......................................................................................................... 64
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Pegasus Separation System Test Unit......................................................................................... 64
10.19.
10.20. Round-the-Clock Payload Support ............................................................................................... 64
10.21. Stage 2 Onboard Camera ............................................................................................................ 64
10.22. Thermal Coated Forward Separation Ring .................................................................................. 64
10.23. 43 cm (17 in.) Payload Adapter .................................................................................................... 65
10.24. Multiple Payload Adapters Including Related Mission Integration Support.................................. 65
10.25. Dual Payload Adapter with 97 cm (38 in.) Primary PA................................................................. 65
10.25.1. Dual Payload Adapter with 59 cm (23 in.) Primary PA ........................................................ 66
10.25.2. Dual Payload Adapter with 43 cm (17 in.) Primary PA ........................................................ 66
10.26. Secondary Payload Adapters for Nonseparating Secondary Payloads....................................... 67
10.26.1. 59 cm (23 in.) or 43 cm (17 in.) PA for Nonseparating Secondary Payloads ...................... 67
10.26.2. Load-Bearing Nonseparating Secondary Payload............................................................... 67
10.26.3. Secondary Payload Adapters for Separating Secondary Payloads..................................... 67
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TABLE OF CONTENTS (CONTINUED)
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10.26.4. 43 cm (17 in.) Payload Adapter............................................................................................ 67
10.26.5. 59 cm (23 in.) Payload Adapter............................................................................................ 67
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LIST OF FIGURES
PAGE
Figure 1-1. Pegasus Rollout ........................................................................................................................ 1
Figure 1-2. Pegasus Launch Locations ....................................................................................................... 2
Figure 2-1. Pegasus XL on the Assembly and Integration Trailer (AIT) ...................................................... 3
Figure 2-2. Expanded View of Pegasus XL Configuration........................................................................... 4
Figure 2-3. Principle Dimensions of Pegasus XL (Reference Only)............................................................ 5
Figure 2-4. Typical Pegasus XL Motor Characteristics in Metric (English) Units ........................................ 6
Figure 2-5. Typical Attitude and Guidance Modes Sequence .....................................................................7
Figure 3-1. Pegasus XL Mission Profile to 741 km (400nmi) Circular, Polar Orbit with a 227 kg
(501 lbm) Payload ..................................................................................................................... 9
Figure 3-2. Pegasus XL with HAPS Mission Profile to 741 km (400nmi) Circular, Polar Orbit with
a 227 kg (501 lbm) Payload..................................................................................................... 10
Figure 3-3. Pegasus XL Without HAPS Performance Capability............................................................... 11
Figure 3-4. Pegasus XL With HAPS Performance Capability.................................................................... 12
Figure 3-5. 3-sigma Injection Accuracies Typical of Pegasus XL Missions............................................... 12
Figure 3-6. Typical and Recent Pegasus Orbital Accuracy ....................................................................... 13
Figure 4-1. Factors of Safety for Payload Design and Test....................................................................... 15
Figure 4-2. Payload Testing Requirements ...............................................................................................15
Figure 4-3. Pegasus Design Limit Load Factors........................................................................................ 16
Figure 4-4. Pegasus XL Maximum Quasi Steady Acceleration as a Function of Payload Weight............ 17
Figure 4-5. Pegasus Net CG Load Factor Predictions .............................................................................. 17
Figure 4-6. Motor Ignition Transient Shock Response Spectrum Specification ........................................ 18
Figure 4-7. Payload Interface Random Vibration Specification ................................................................. 18
Figure 4-8. Shock Environment at Base of the Payload ............................................................................ 19
Figure 4-9. Payload Acoustic Environment................................................................................................ 19
Figure 4-10. Representative Fairing Internal Pressure Profile During Captive Carry................................ 20
Figure 4-11. Representative Fairing Internal Pressure Profile During Powered Flight.............................. 20
Figure 4-12. Pegasus XL RF Emitters and Receivers ............................................................................... 22
Figure 4-13. Carrier Aircraft RF Emitters and Receivers ........................................................................... 22
Figure 5-1. Payload Fairing Static Envelope with 97 cm (38 in.) Diameter Payload ................................. 25
Figure 5-2. Payload Fairing Dynamic Envelope with 97 cm (38 in.) Diameter Payload ............................ 26
Figure 5-3. Payload Fairing Static Envelope with 59 cm (23 in.) Diameter Payload ................................. 27
Figure 5-4. Payload Fairing Dynamic Envelope with 59 cm (23 in.) Diameter Payload ............................ 28
Figure 5-5. Payload Fairing Access Door Placement Zones (shown with optional second door shown) . 29
Figure 5-6. Nonseparable Payload Mechanical Interface..........................................................................30
Figure 5-7. 97 cm (38 in.) Separable Payload Interface ............................................................................ 31
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LIST OF FIGURES (CONTINUED)
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Figure 5-8. 59 cm (23 in.) Separable Payload Interface ............................................................................ 32
Figure 5-9. Payload Separation Velocities Using the Standard Separation System ................................. 33
Figure 5-10. Standard Payload Electrical Interface ................................................................................... 34
Figure 5-11. Payload Mass vs. c.g. Location on X Axis............................................................................. 36
Figure 5-12. Payload Mass Property Measurement Error Tolerances ...................................................... 37
Figure 5-13. Detailed RCS Dead Band Zone............................................................................................. 37
Figure 5-14. Pegasus/OCA Interface Details............................................................................................. 38
Figure 6-1. Mission Integration Management Structure............................................................................. 40
Figure 6-2. Summary of Typical Working Groups...................................................................................... 42
Figure 6-3. Typical Mission Cycle .............................................................................................................. 44
Figure 6-4. Applicable Safety Requirements ............................................................................................. 44
Figure 6-5. Safety Approval Process ......................................................................................................... 46
Figure 7-1. Typical Processing Flow.......................................................................................................... 47
Figure 7-2. Typical Pegasus Integration and Test Schedule ..................................................................... 48
Figure 7-3. Orbital Carrier Aircraft Hot Pad Area at VAFB ........................................................................ 48
Figure 7-4. Pegasus Integration................................................................................................................. 49
Figure 7-5. Typical Pegasus Launch Checklist Flow ................................................................................. 52
Figure 8-1. Documentation Produced by Orbital for Commercial Pegasus Launch Services ................... 53
Figure 8-2. Documentation Required by Orbital for Commercial Pegasus Launch Services.................... 54
Figure 9-1. Load-Bearing Spacecraft Configuration .................................................................................. 57
Figure 9-2. Dual Payload Attach Fitting Configuration............................................................................... 58
Figure 10-1. Hydrazine Auxillary Propulsion System (HAPS) ................................................................... 62
LIST OF APPENDICES
A. PAYLOAD QUESTIONNAIRE..............................................................................................................A-1
B. VAFB VEHICLE ASSEMBLY BUILDING CAPABILITIES ....................................................................B-1
C. LAUNCH RANGE INFORMATION...................................................................................................... C-1
D. PEGASUS FLIGHT HISTORY ............................................................................................................ D-1
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LIST OF ACRONYMS
3DOF Three Degrees of Freedom
6DOF Six Degrees of Freedom
A, Amps Amperes
AACS Airborne Air Conditioning System
ac Alternating Current
A/C Air Conditioning
AFB Air Force Base
AIT Assembly and Integration Trailer
ARAR Accident Risk Asessment Report
ARO After Receipt of Order
ASE Airborne Support Equipment
ATP Authority to Proceed
AWG American Wire Gauge
C/CAM Collision/Contamination Avoidance
Maneuver
C Centigrade
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
DoD Department of Defense
DoT Department of Transportation
DPA Dual Payload Adapter
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
FAS Fin Actuation System
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
GOP Ground Operations Plan
GPS Global Positioning System (NAVSTAR)
Grms Gravity Root Mean Squared
GSE Ground Support Equipment
H/W Hardware
h Height
HAPS Hydrazine Auxiliary Propulsion System
HEPA High Efficiency Particulate Air
HF High Frequency
HVAC Heating, Ventilating, and Air
Conditioning
Hz Hertz
ICD Interface Control Document
IEEE Institute of Electrical and Electronic
Engineers
ILC Initial Launch Capability
IMU Inertial Measurement Unit
in. Inch
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/s Meters Per Second
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
MPS Mission Planning Schedule
MRR Mission Readiness Review
ms Millisecond
MSD Mission Specification Document
MSPSP Missile System Prelaunch Safety
Package
MUX Multiplexer
N2 Nitrogen
N/A Not Applicable
N Newtons
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NASA National Aeronautics and Space
Administration
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
P/L Payload
PA Payload Adapter
PDR Preliminary Design Review
PDU Pyrotechnic Driver Unit
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
psig Pounds per Square Inch Gauge
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
Pegasus User’s Guide
rpm Revolutions Per Minute
RTB Return to Base
RSS Root Summed Squared
RTV Room Temperature Vulcanizing
S&A Safe & Arm
S/N Serial Number
S/W Software
scfm Standard Cubic Feet Per Minute
sec Second(s)
SSPP System Safety Program Plan
SWC Soft Wall Cleanroom
TLM Telemetry
T.O. Take-Off
TPS Thermal Protection System
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
VT Verification Test
WFF Wallops Flight Facility
WR Western Range (USAF)
XL Extended Length (Pegasus)
YFS Yield Factor of Safety
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1. INTRODUCTION
On August 10, 1989, Orbital Sciences Corporation
(Orbital) rolled out the first commercially
developed space launch vehicle for providing
satellites to low earth orbit (see Figure 1-1). Over
the past 21 years, the “winged rocket” known as
Pegasus has proven to be the most successful in
its class, placing over 78 satellites in orbit with 40
launches as of April 2010.
Figure 1-1. Pegasus Rollout
This Pegasus User’s Guide is intended to
familiarize mission planners with the capabilities
and services provided with a Pegasus launch.
The Pegasus XL was developed as an increased
performance design evolution from the original
Pegasus vehicle to support NASA and the USAF
performance requirements, and is now the
baseline configuration for all commercial Pegasus
launches.
Pegasus is a mature and flight-proven launch
system that has demonstrated consistent accuracy
and dependable performance. The Pegasus
launch system has achieved a high degree of
reliability through its significant flight experience.
Pegasus offers a variety of capabilities that are
uniquely suited to small spacecraft. These
capabilities and features provide the small
Pegasus User’s Guide
spacecraft customer with greater mission utility in
the form of:
A range of custom payload interfaces and
services to accommodate unique small
spacecraft missions;
Payload support services at the Pegasus
Vehicle Assembly Building (VAB) at
Vandenberg Air Force Base (VAFB),
California;
Horizontal payload integration;
Shared payload launch accommodations for
more cost-effective access to space as
compared to Dual Launches;
Portable air-launch capability from worldwide
locations to satisfy unique mission
requirements; and
Fast, cost-effective, and reliable access to
space.
The mobile nature of Pegasus allows Orbital to
integrate the spacecraft to the Pegasus XL in our
integration facility, the VAB, and ferry the launchready system to a variety of launch ranges.
Pegasus has launched from a number of launch
locations worldwide (see Figure 1-2).
The unique mobile capability of the Pegasus
launch system provides flexibility and versatility to
the payload customer. The Pegasus launch
vehicle can accommodate integration of the
spacecraft at a customer desired location, as well
as optimize desired orbit requirements based on
the initial launch location. In 1997, after final build
up of the rocket at the VAB, Pegasus was mated
to the Orbital Carrier Aircraft (OCA) and ferried to
Madrid, Spain, to integrate Spain’s MINISAT-01
satellite. Following integration of the satellite,
Pegasus was then ferried to the island of Gran
Canaria for launch. The successful launch of
Spain’s MINISAT-01 satellite demonstrated
Pegasus’ ability to accommodate the payload
provider’s processing and launch requirements at
locations better suited to the customer rather than
the launch vehicle. This unprecedented launch
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Pegasus User’s Guide
Figure 1-2. Pegasus Launch Locations
vehicle approach is an example of the Pegasus
way of providing customer oriented launch service.
In the interest of continued process improvement
and customer satisfaction, the Pegasus Program
successfully completed a 1-year effort of ISO 9001
certification. In July 1998, Orbital’s Launch
Systems Group was awarded this internationally
recognized industry benchmark for operating a
quality management system producing a
qualityproduct and service. Since that time,
Orbital has achieved third party certification to
ISO9001:2008 and AS9100B, providing even
greater assurance of mission success. In addition
to our AS9100B certification, NASA has granted
the Pegasus XL Launch Vehicle a Category 3
certification that qualifies Pegasus to launch
NASA’s highest value spacecraft.
Pegasus is a customer oriented and responsive
launch vehicle system. From Pegasus’ com-
mercial heritage comes the desire to continually
address the payload customer market to best
accommodate its needs. The Pegasus launch
vehicle system has continually matured and
evolved over its 21-year history. This ability and
desire to react to the customer has produced the
single most successful launch vehicle in its class.
To ensure our goal of complete customer
satisfaction, a team of managers and engineers is
assigned to each mission from “contract award to
post-flight report.” This dedicated team is
committed to providing the payload customer
100% satisfaction of mission requirements.
Each Pegasus mission is assigned a mission team
led by a Mission Manager and a Mission Engineer.
The mission team is responsible for mission
planning and scheduling, launch vehicle
production coordination, payload integration
services, systems engineering, mission-peculiar
design and analysis, payload interface definition,
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range coordination, launch site processing, and
operations. The mission team is responsible for
ensuring all mission requirements have been
satisfied.
2. PEGASUS DESCRIPTION
The Pegasus User’s Guide is dedicated to the
discussion of the Pegasus XL configuration,
capabilities, and associated services.
2.1. Pegasus XL Vehicle Description
Pegasus XL is a winged, three-stage, solid rocket
booster that weighs approximately 23,130 kg
(51,000 lbm), and measures 16.9 m (55.4 ft) in
length and 1.27 m (50 in.) in diameter, and has a
wing span of 6.7 m (22 ft). Figure 2-1 shows the
Pegasus on the Assembly Integration Trailer (AIT).
Pegasus is lifted by the OCA to a level flight
condition of about 11,900 m (39,000 ft) and Mach
0.82. Five seconds after release from the OCA
Stage 1 motor ignition occurs. The vehicle’s
autonomous guidance and flight control system
provide the guidance necessary to insert payloads
into a wide range of orbits.
Figure 2-1. Pegasus XL on the Assembly and
Integration Trailer (AIT)
Figure 2-2 shows an expanded view of the
Pegasus XL configuration. The Pegasus Vehicle
design combines flight-proven technologies, and
conservative design margins to achieve
performance and reliability. The vehicle
incorporates eight major elements:
Pegasus User’s Guide
Three solid rocket motors;
A payload fairing;
An avionics assembly;
A lifting wing;
Aft skirt assembly including three movable
control fins; and
A payload interface system.
Pegasus also has an option for a liquid propellant
fourth stage, HAPS (see Section 10). Figure 2-3
illustrates Pegasus XL’s principle dimensions.
2.1.1. Solid Rocket Motors
The three solid rocket motors were designed and
optimized specifically for Pegasus and include
features that emphasize reliability and
manufacturability. The design was developed
using previously flight-proven and qualified
materials and components. Common design
features, materials, and production techniques are
applied to all three motors to maximize cost
efficiency and reliability. These motors are fully
flight-qualified. Typical motor characteristics are
shown in Figure 2-4.
2.1.2. Payload Fairing
The Pegasus payload fairing consists of two
composite shell halves, a nose cap integral to a
shell half, and a separation system. Each shell
half is composed of a cylinder and ogive sections.
The two halves are held together with a base
frangible joint, two titanium straps along the
cylinder and a retention bolt in the nose. A cork
and Room Temperature Vulcanizing (RTV)
Thermal Protection System (TPS) provides
protection to the graphite composite fairing
structure. The amount of TPS applied has been
determined to optimize fairing performance and
payload environmental protection.
The two straps are tensioned using bolts, which
are severed during fairing separation with
pyrotechnic bolt cutters, while the retention bolt in
the nose is released with a pyrotechnic separation
nut. The base of the fairing is separated with
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Pegasus User’s Guide
Figure 2-2. Expanded View of Pegasus XL Configuration
Orbital’s low-contamination frangible separation
joint. These ordnance events are sequenced for
proper separation dynamics. A hot gas generator
internal to the fairing is also activated at
separation to pressurize two piston-driven pushoff
thrusters. These units, in conjunction with cams,
force the two fairing halves apart. The halves
rotate about fall-away hinges, which guide them
away from the satellite and launch vehicle.
The fairing and separation system were fully
qualified through a series of structural, functional,
and contamination ground vacuum tests and have
been successfully flown on all Pegasus XL
missions. Section 5 presents a more detailed
description of the fairing separation sequence and
the satellite dynamic envelope.
2.1.3. Avionics
The Pegasus avionics system is a digital
distributed processor design that implements
developments in hardware, software, communi-
cations, and systems design. Mission reliability is
achieved by the use of simple designs, high
reliability components, high design margins, and
extensive testing at the component, subsystem,
and system level.
The heart of the Pegasus avionics system is a
multiprocessor, 32-bit flight computer. The flight
computer communicates with the Inertial
Measurement Unit (IMU), the launch panel
electronics on the carrier aircraft, and all vehicle
subsystems using standard RS-422 digital serial
data links. Most avionics on the vehicle feature
integral microprocessors to perform local
processing and to handle communications with the
flight computer. This RS-422 architecture is
central to Pegasus rapid integration and test, as it
allows unit and system-level testing to be
accomplished using commercially available
ground support equipment with off-the-shelf
hardware.
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Figure 2-3. Principle Dimensions of Pegasus XL (Reference Only)
2.1.4. Flight Termination System
The Pegasus Flight Termination System (FTS)
supports ground-initiated command destruct as
well as the capability to sense inadvertent stage
separation and automatically destruct the rocket.
The FTS is redundant, with two independent safe
and arm devices, receivers, logic units, and
batteries.
2.1.5. Attitude Control Systems
After release from the OCA, the Pegasus attitude
control system is fully autonomous. A combination
of open-loop steering and closed-loop guidance is
employed during the flight. Stage 1 guidance
utilizes a pitch profile optimized by ground
simulations. Stage 2 and Stage 3 guidance uses
an adaptation of an algorithm that was first
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Figure 2-4. Typical Pegasus XL Motor Characteristics in Metric (English) Units
developed for the Space Shuttle ascent guidance.
Attitude control is closed-loop.
The vehicle attitude is controlled by the Fin
Actuator System (FAS) during Stage 1 flight. This
consists of electrically actuated fins located at the
aft end of Stage 1. For Stage 2 and Stage 3 flight,
a combination of electrically activated Thrust
Vector Controllers (TVCs) on the Stage 2 and
Stage 3 solid motor nozzles and a GN2 Reaction
Control System (RCS) located on the avionics
section, control the vehicle attitude.
Figure 2-5 summarizes the attitude and guidance
modes during a typical flight, although the exact
sequence is controlled by the Mission Data Load
(MDL) software and depends on mission-specific
requirements.
2.1.6. Telemetry Subsystem
The Pegasus XL telemetry system provides realtime health and status data of the vehicle avionics
system, as well as key information regarding the
position, performance, and environment of the
Pegasus XL vehicle. This data is used by Orbital
and the range safety personnel to evaluate system
performance.
Pegasus contains two separate telemetry
systems. The first provides digital data through
telemetry multiplexers (MUXs), which gather data
from each sensor, digitize it, then relay the
information to the flight computer. This Pegasus
telemetry stream provides data during ground
processing, checkout, captive carry, and during
launch. During captive carry, Pegasus telemetry
is downlinked to the ground and recorded onboard
the OCA. Some payload telemetry data can be
interleaved with Pegasus data as a nonstandard
service. The second system provides analog
environments data, which are transmitted via a
wideband data link and recorded for post-flight
evaluation.
2.1.7. Major Structural Subsystems
2.1.7.1. Wing
The Pegasus wing uses a truncated delta platform
with a double wedge profile. Wing panels are
made of a graphite-faced foam sandwich.
Channel section graphite spars carry the primary
bending loads and half-ribs, and reinforcing layups further stabilize the panels and reduce stress
concentrations. The wing central box structure
has fittings at each corner that provide the
structural interface between the Pegasus and the
OCA.
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Figure 2-5. Typical Attitude and Guidance Modes Sequence
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 aft
skirt is an all-aluminum structure of conventional
ring and stressed-skin design with machined
bridge fittings for installation of the
electromechanical fin actuators. The skirt is
segmented to allow installation around the first
stage nozzle. Fin construction is a one-piece solid
foam core and wet-laid graphite composite
construction around a central titanium shaft.
2.1.7.3. Payload Interface Systems
Multiple mechanical and electrical interface
systems currently exist to accommodate a variety
of spacecraft designs. Section 5.0 describes
these interface systems. To ensure optimization
of spacecraft requirements, payload specific
mechanical and electrical interface systems can
be provided to the payload customer. Payload
mechanical fit checks and electrical interface
testing with these spacecraft unique interface
systems are encouraged to ensure all spacecraft
requirements are satisfied early in the processing
flow.
2.2. Orbital Carrier Aircraft
Orbital furnishes and operates the OCA. After
integration at Orbital’s West Coast integration site
at VAFB, the OCA can provide polar and high-
inclination launches utilizing the tracking,
telemetry, and command (TT&C) facilities of the
WR. The OCA can provide lower inclination
missions from the East Coast using either the
NASA or ER TT&C facilities or from the Reagan
Test Site from the Kwajalein Atoll, as well as
equatorial missions from the Kwajalein Atoll. The
OCA is made available for mission support on a
priority basis during the contract-specified launch
window.
The unique OCA-Pegasus launch system
accommodates two distinctly different launch
processing and operations approaches for nonVAFB launches. One approach (used by the
majority of payload customers) is to integrate the
Pegasus and payload at the VAB and then ferry
the integrated Pegasus and payload to another
location for launch. This approach is referred to
as a “ferry mission.” The second approach is
referred to as a “campaign mission.” A campaign
mission starts with the build up of the Pegasus at
the VAB. The Pegasus is then mated to the OCA
at VAFB and ferried to the integration site where
the Pegasus and payload are fully integrated and
tested. At this point, the launch may either occur
at the integration site, or the integrated Pegasus
and payload may be ferried to another location for
launch.
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The OCA also has the capability to ferry Pegasus
across the United States or across the ocean
(depending on landing site) to support ferry and
campaign missions.
3. GENERAL PERFORMANCE CAPABILITY
This section describes the orbital performance
capabilities of the Pegasus XL vehicle with and
without the optional HAPS described in Section
10. Together these configurations can deliver
payloads to a wide variety of circular and elliptical
orbits and trajectories, and attain a complete
range of prograde and retrograde inclinations
through a suitable choice of launch points and
azimuths. In general, the optional HAPS will
provide additional performance at higher altitudes,
as well as providing a more accurate insertion
orbit capability.
From the WR, Pegasus can achieve inclinations
between 70° and 130°. A broader range of
inclinations may be achievable, subject to
additional analyses and coordination with Range
authorities. Additionally, lower inclinations can be
achieved through dog-leg trajectories, with a
commensurate reduction in performance. Some
specific inclinations within this range may be
limited by stage impact point or other restrictions.
Other inclinations can be supported through use of
Wallops Flight Facility (WFF), Eastern Range
(ER), Reagan Test Site (RTS) Kwajalein, or other
remote TT&C sites. Pegasus requirements for
remote sites are listed in Appendix D.
3.1. Mission Profiles
This section describes circular low earth orbit
mission profiles. Performance quotes for noncircular orbits will be provided on a missionspecific basis.
Profiles of typical missions performed by Pegasus
XL with and without HAPS are illustrated in
Figure 3-1 and Figure 3-2. The depicted profile
begins after the OCA has reached the launch
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point, and continues through orbit insertion. The
time, altitude, and velocity for the major ignition,
separation, and burnout events are shown for a
typical trajectory that achieves a 741 km (400 nm)
circular, polar (90° inclination) orbit after launch
from the WR. These events will vary based on
mission requirements.
The typical launch sequence begins with release
of Pegasus from the carrier aircraft at an altitude
of approximately 11,900 m (39,000 ft) and a speed
of Mach 0.82. Approximately 5 seconds after
drop, when Pegasus has cleared the aircraft,
Stage 1 ignition occurs. The vehicle quickly
accelerates to supersonic speed while beginning a
pull up maneuver. Maximum dynamic pressure is
experienced approximately 30 seconds after
ignition. At approximately 15-20 seconds, a
maneuver is initiated to depress the trajectory and
the vehicle transitions to progressively lower
angles of attack.
Stage 2 ignition occurs shortly after Stage 1
burnout, and the payload fairing is jettisoned
during Stage 2 burn as quickly as fairing dynamic
pressure and payload aerodynamic heating
limitations will allow, approximately 112,000 m
(366,000 ft) and 121 seconds after drop. Stage 2
burnout is followed by a long coast, during which
the payload and Stage 3 achieve orbital altitude.
For a non-HAPS Pegasus configuration, Stage 3
then provides the additional velocity necessary to
circularize the orbit. Stage 3 burnout typically
occurs approximately 10 minutes after launch and
2,200 km (1,200 nm) downrange of the launch
point.
An FAS, in conjunction with three aerodynamic
fins, provides attitude control from drop through
Stage 1 separation. Pitch and yaw attitude control
during Stage 2 and Stage 3 powered flight is
provided by the motor TVC system while roll
attitude is controlled by the nitrogen cold gas RCS.
The RCS also provides three-axis control during
coast phases of the trajectory.
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Figure 3-1. Pegasus XL Mission Profile to 741 km (400nmi) Circular, Polar Orbit with a 227 kg
(501 lbm) Payload
3.2. Performance Capability
Performance capabilities to various orbits for the
Pegasus XL are illustrated in Figure 3-3 and
Figure 3-4 (HAPS configuration). These
performance data were generated using the
Program to Optimize Simulated Trajectories
(POST), which is described below. Precise
performance capabilities to specific orbits are
typically provided per the documentation schedule
shown in Section 8.0.
3.3. Trajectory Design Optimization
Orbital designs a unique mission trajectory for
each Pegasus flight to maximize payload
performance while complying with any applicable
payload and launch vehicle constraints. In this
process, a 3DOF simulation is developed using
the current Pegasus mass properties,
aerodynamic models, and motor ballistics data,
and the desired target orbit and any applicable
trajectory constraints are specified. POST then
uses a set of specified control parameters to
iterate on the trajectory design until an optimal
solution is identified which maximizes performance
to the desired target orbit subject to the specified
constraints. Typically, these constraints may
include limitations on the angle of attack profile,
dynamic loading constraints, payload
environmental constraints such as heat rate, and
Range-imposed constraints on the launch azimuth
and spent stage impact locations. After POST has
been used to determine the optimal trajectory
design, a high-fidelity, Pegasus-specific, 6DOF
simulation is then developed to conduct detailed
trajectory analyses to verify the acceptability of the
trajectory design and to verify robust control
system stability margins.
3.4. Orbit Insertion Accuracy
The estimated orbit insertion errors for Pegasus
vary from mission to mission and are influenced by
a variety of factors including the target orbit,
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Figure 3-2. Pegasus XL with HAPS Mission Profile to 741 km (400nmi) Circular, Polar Orbit with a
227 kg (501 lbm) Payload
trajectory design, payload mass, and the guidance
strategy requested by the payload. As a result,
the specific Pegasus orbit accuracy capabilities for
a particular mission are generally determined only
after these mission-specific details are defined and
detailed mission-specific analyses have been
performed. However, Figure 3-5 provides
estimates of 3-sigma orbit insertion errors for both
Pegasus XL and Pegasus XL with HAPS vehicle
configurations, which are representative of typical
Pegasus missions. For non-HAPS configurations,
these errors are generally dominated by the
impulse variability associated with Stage 3. This
variability is also responsible for the generally
larger magnitude errors for the non-insertion apse
relative to the insertion apse.
3.4.1. Actual Pegasus Insertion Accuracies
Figure 3-6 shows the actual Pegasus orbital
insertion accuracies achieved for all missions
since Flight 10. As this figure demonstrates, a
large majority of these missions resulted in
perigee and apogee altitudes within 30 km of the
desired target values and inclination errors of less
than 0.05 degrees.
3.4.2. Error-Minimizing Guidance Strategies
Due to the large amount of actual flight experience
Pegasus has accumulated to date, the Pegasus
Program has been able to continually refine and
improve the fidelity and accuracy of the Pegasus
vehicle simulation. This process has allowed us to
develop a high degree of confidence in the
Pegasus simulation analysis results and to
accurately predict mission performance in flight.
To ensure that even a 3-sigma low-performing
Pegasus vehicle will achieve the required orbit,
Pegasus trajectories include a 67 m/sec
(220 ft/sec) guidance reserve. Pegasus flight
software provides the capability to manage this
reserve through the use of a variety of different
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Figure 3-3. Pegasus XL Without HAPS Performance Capability
guidance strategies that are designed and tailored
to meet specific mission objectives. These
strategies fall into several basic categories:
(1) Minimize Insertion Errors. Using this strategy,
the guidance system manages the excess
vehicle energy by implementing out-of-plane
turning during Stage 2 and Stage 3 burns as
required, and by adjusting the timing of
Stage 3 ignition. This “energy-scrubbing”
strategy results in the smallest possible
insertion errors for both apogee and perigee
altitudes.
(2) Maximize Insertion Altitude. Using this
strategy, excess vehicle performance is
conserved to maximize the altitude at
insertion. This allows the customer to achieve
the highest possible circular orbit altitude
based on the actual vehicle performance while
minimizing the eccentricity of the final orbit.
(3) Maximize Insertion Velocity. Using this
strategy, excess vehicle performance is
conserved to maximize velocity at insertion.
This allows the customer to use the excess
guidance reserve to increase the expected
apogee (non-insertion apse) altitude while
continuing to maintain a precise perigee
(insertion apse) altitude.
(4) Some Combination of (2) and (3). Options 2
and 3 represent the two endpoints of a
spectrum of potential guidance strategies that
can be combined and tailored to achieve
mission-specific guidance objectives. Both
insertion altitude and velocity may be
maximized to achieve the highest possible
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Figure 3-4. Pegasus XL With HAPS Performance Capability
orbit energy, or specific altitude and velocity
thresholds may be defined, which trigger
energy-scrubbing only in the event that the
thresholds are exceeded. The optimal
strategy for a particular mission will therefore
depend on the specific guidance objectives.
Figure 3-5. 3-sigma Injection Accuracies
Typical of Pegasus XL Missions
3.5. Collision/Contamination Avoidance
Maneuver
Following orbit insertion, the Pegasus Stage 3
RCS or HAPS will perform a Collision/Contam-
ination Avoidance Maneuver (C/CAM). The
C/CAM consists of a series of maneuvers
designed to both minimize payload contamination
and the potential for recontact between Pegasus
hardware and the separated payload.
Orbital will perform a recontact analysis for post
separation events. Orbital and the payload
contractor are jointly responsible for determination
of whether a C/CAM is required.
A typical C/CAM (for a non-HAPS configuration)
consists of the following steps:
(1) At payload separation +3 seconds, the launch
vehicle performs a 90° yaw maneuver
designed to direct any remaining Stage 3
motor impulse in a direction that will increase
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Figure 3-6. Typical and Recent Pegasus Orbital Accuracy
the separation distance between the two
bodies.
(2) At payload separation +300 seconds, the
launch vehicle begins a “crab-walk” maneuver.
This maneuver, performed through a series of
RCS thruster firings, is designed to impart a
small amount of delta velocity in a direction
designed to increase the separation distance
between Pegasus and the payload. The
maneuver is terminated approximately
600 seconds after separation.
At the completion of the C/CAM, all remaining
nitrogen and/or hydrazine is depleted.
4. PAYLOAD ENVIRONMENTS
The following subsections present the maximum
payload environment levels during Pegasus
captive carry and powered flight. The acoustic,
vibration, shock, and acceleration environments
presented below apply to the launch vehicle with a
single payload using either the 38" or 23" payload
adapter. The payload environments associated
with the use of alternative separation systems, a
nonseparating payload interface, or multiple
payload attach fittings will differ from those
presented below.
The electromagnetic radiation and thermal
environments presented below apply to all launch
vehicle and payload configurations.
4.1. Design Loads
The primary support structure for the spacecraft
shall possess sufficient strength, rigidity, and other
characteristics required to survive the critical
loading conditions that exist within the envelope of
handling and mission requirements, including
worst-case predicted ground, flight, and orbital
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loads. It shall survive those conditions in a
manner that ensures safety and that does not
reduce the mission success probability. The
primary support structure of the spacecraft shall
be electrically conductive to establish a single
point electrical ground. Spacecraft design loads
are defined as follows:
Design Limit Load — The maximum predicted
ground-based, captive carry, or powered flight
load, including all uncertainties.
Design Yield Load — The Design Limit Load
multiplied by the required Yield Factor of
Safety (YFS) indicated in Figure 4-1. The
payload structure must have sufficient strength
to withstand simultaneously the yield loads,
applied temperature, and other accompanying
environmental phenomena for each design
condition without experiencing detrimental
yielding or permanent deformation.
Design Ultimate Load — The Design Limit
Load multiplied by the required Ultimate
Factor of Safety (UFS) indicated in Figure 4-1.
The payload structure must have sufficient
strength to withstand simultaneously the
ultimate loads, applied temperature, and other
accompanying environmental phenomena
without experiencing any fracture or other
failure mode of the structure.
4.2. Payload Testing and Analysis
Sufficient payload testing and/or analysis must be
performed to ensure the safety of ground and
aircraft crews and to ensure mission success. The
payload design must comply with the testing and
design factors of safety in Figure 4-1 and the FAA
regulations for the carrier aircraft listed in the
CFR14 document, FAR Part 25. UFS shown in
Figure 4-1 must be maintained per Orbital SSD
TD-0005. At a minimum, the following tests must
be performed:
Structural Integrity — Static loads or other
tests shall be performed that combine to
encompass the acceleration load environment
presented in Section 4.3. Test level
requirements are defined in Figure 4-1.
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Random Vibration — Test level requirements
are defined in Figure 4-2.
4.3. Payload Acceleration Environment
Maximum expected loads during captive carry and
launch are shown in Figures 4-3, 4-4, and 4-5.
The Pegasus air-launch operation results in a
launch vehicle/OCA separation transient at drop.
The drop transient acceleration limits presented
here are based on two assumptions:
(1) Pegasus Standard 23” or 38” payload
separation system is used.
(2) The first fundamental lateral frequency of the
spacecraft cantilevered at the payload
interface (excluding the payload separation
system) is greater than 20 Hz.
If either assumption is violated, mission-specific
analyses are required. For all missions, accurate
estimation of the drop transient loading requires a
coupled loads analysis (CLA), which uses Orbital
and customer-provided finite element models to
predict the transient environment (see Section
8.3.3 for details).
Transient loading also exists due to motor ignition.
Stage 1 provides the worst-case loading due to
motor ignition. The Stage 1 ignition acceleration
limits at the payload interface are listed in
Figure 4-3. The Stage 1 shock response
spectrum (SRS) at the payload interface is shown
in Figure 4-6. As is the case with the drop
transient, accurate estimation of loading requires a
CLA. The Stage 1 ignition transient CLA requires
finite element models of the Pegasus avionics
structure, payload separation system, and the
payload.
4.4. Payload Random Vibration Environment
The maximum expected random vibration levels at
the payload interface are shown in Figure 4-7.
Random vibration data recorded during multiple
Pegasus missions was used to create this overall
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Figure 4-1. Factors of Safety for Payload Design and Test
envelope that encompasses all phases of a
Pegasus launch operation including OCA takeoff,
captive carry, and powered flight.
A +3 dB factor should be added to this spectrum
for 75 seconds in each axis for payload standard
vibration testing to account for fatigue duration
effects to encompass at least two launch attempts
and powered flight.
4.5. Sinusoidal Vibration
The Pegasus launch vehicle has no significant
sustained sinusoidal vibration environments during
captive carry or powered flight.
4.6. Payload Shock Environment
The maximum expected shock response spectrum
at the base of the payload from all launch vehicle
events is shown in Figure 4-8. The flight limit
levels are derived from ground stage and payload
separation test data assuming a 38” Orbitalsupplied separation system.
Figure 4-2. Payload Testing Requirements
4.7. Payload Acoustic Environment
The maximum expected acoustic levels within the
payload fairing are shown in Figure 4-9. Acoustic
data recorded during previous Pegasus missions
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was used to create this overall envelope that
encompasses all phases of Pegasus launch
operation including OCA takeoff, captive carry,
and powered flight.
A +6 dB factor should be added to this spectrum
for 75 seconds for payload standard acoustic
testing to account for fatigue duration effects to
encompass at least two launch attempts and
powered flight.
4.8. Pressure Profile
Due to the low pressure decay rate associated
with OCA ascent and low initial static pressure at
drop, the depressurization rates for the Pegasus
payload fairing are less than 0.3 psi/sec. The
internal pressure at fairing jettison is well below
0.1 psia. Representative pressure profiles for
captive carry and powered flight are provided in
Figures 4-10 and 4-11.
4.9. Payload Thermal Environment
The payload thermal environment is maintained
during all phases of integrated operations
including payload processing, fairing
encapsulation, transportation of the launch
vehicle, ground operations at the flight line and
launch operations.
4.9.1. Payload Processing
During payload processing, the temperature and
humidity of the spacecraft processing areas within
Building 1555 are maintained within a range of 18
to 29 ºC (64.4 to 84.2 ºF) and ≤55%, respectively.
Following encapsulation of the payload, but prior
to transportation of the Pegasus vehicle to the Hot
Pad, the fairing is continuously purged with filtered
air. The temperature and humidity limits are the
same as listed above. The flowrate of air through
the fairing is maintained between 50 and 200 cfm.
The air flow enters the fairing forward of the
payload and exits aft of the payload. There are
baffles on the inlet that minimize the impingement
velocity of the air on the payload.
4.9.2. Transportation
During transportation of the Pegasus vehicle to the
Hot Pad, the fairing is continuously purged with
filtered and dried ambient air. The air temperature
is not actively controlled; however, transportation
operations are performed only when the ambient
temperature ensures that the air supplied to the
fairing will be between 2 to 29 ºC (35.6 to 84.2 ºF).
The relative humidity of the air supplied to the
Figure 4-3. Pegasus Design Limit Load Factors
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Figure 4-4. Pegasus XL Maximum Quasi Steady Acceleration as a Function of Payload Weight
Figure 4-5. Pegasus Net CG Load Factor Predictions
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