DELTA II Payload Planners Guide

The Boeing Company Space and Communications Group APRIL 1996 MDC H3224D

DELTA II PAYLOAD PLANNERS GUIDE

The Delta II Payload Planners Guide has been cleared for public release by the Chief—Air Force Division, Directorate for Freedom of Information and Security Review, Office of the Assistant Secretary of Defense, as stated in letter______,dated95-S-3826/L ______1995.3 October

THIS DOCUMENT SUPERSEDES PREVIOUS ISSUES OF THE COMMERCIAL DELTA II PAYLOAD PLANNERS GUIDE, MDC H3224C, DATED OCTOBER 1993

The Boeing Company 5301 Bolsa Avenue, Huntington Beach, CA 92647-2099 (714) 896-3311 02715REU9.1 PUBLICATION NOTICE TO HOLDERS OF THE DELTA II PAYLOAD PLANNERS GUIDE

The Delta II Payload Planners Guide will be revised periodically to incorporate the latest information. You are encouraged to return the Revision Service Card below to ensure that you are included on the mailing list for future revisions of the Delta II Payload Planners Guide. Changes to your address should be noted in the space provided.

Please forward any comments or suggestions you have concerning content or format. Inquiries to clarify or interpret this material should be directed to:

Delta Launch Services c/o The Boeing Company 5301 Bolsa Avenue, (MC H014-C426) Huntington Beach, CA 92647-2099 E-mail: [email protected]

MDC H3224D October 1999 REVISION SERVICE CARD DELTA II PAYLOAD PLANNERS GUIDE CURRENT ADDRESS Check all that apply: Name: Send hardcopy of next revision Title: Send CD-ROM of next revision Department: Address change Mail Stop: Telephone: Customer Comments: Fax: E-mail: Company Name: Address: City: State: Zip Code: Country: Date: 02717REU9.1 PUBLICATION NOTICE TO HOLDERS OF THE DELTA III PAYLOAD PLANNERS GUIDE

The Delta III Payload Planners Guide will be revised periodically to incorporate the latest information. You are encouraged to return the Revision Service Card below to ensure that you are included on the mailing list for future revisions of the Delta III Payload Planners Guide. Changes to your address should be noted in the space provided.

Please forward any comments or suggestions you have concerning content or format. Inquiries to clarify or interpret this material should be directed to:

Delta Launch Services c/o The Boeing Company 5301 Bolsa Avenue, (MC H014-C426) Huntington Beach, CA 92647-2099 E-mail: [email protected]

MDC 99H0068 October 1999 REVISION SERVICE CARD DELTA III PAYLOAD PLANNERS GUIDE NO POSTAGE CURRENT ADDRESS NECESSARY Check all that apply: Name: IF MAILED IN THE Send hardcopy of next revision Title: UNITED STATES Send CD-ROM of next revision Department: Mail Stop: BUSINESS REPLY MAIL Address change FIRST CLASS PERMIT NO. 41, HUNTINGTON BEACH, CA Telephone: Customer Comments: Fax: POSTAGE WILL BE PAID BY ADDRESSEE E-mail: Company Name: Address: Delta Launch Services City: c/o The Boeing Company 5301 Bolsa Avenue, MC H014-C426 State: Huntington Beach, CA 92647-2099 Zip Code: Country: Date: PREFACE

The Delta II Payload Planners Guide (PPG) is issued to potential satellite system users, operators, and spacecraft contractors to provide information regarding the Delta II launch vehicle and its related systems and launch services. On 1 August 1997, the McDonnell Douglas Corporation (MDC) was merged with the Boeing Company. This PPG contains historical references to MDC, which is now the Boeing Company. This document supersedes previous issues of the Delta II Spacecraft Users Manual, MDC H3224, dated July 1987, MDC H3224A/B, dated December 1989, and MDC 93H3224C, dated October 1993. This PPG has been revised to incorporate the latest Delta II upgrades. These improvements continue the Delta II tradition of continuous evolution to provide The Boeing Company customers with increased payload capacity while concurrently improving operability and producibility. Among the Delta II improvements included in this edition, customers will find information on: ■ the latest avionics upgrades – redundant inertial flight control assembly (RIFCA) ■ increased performance from extended nozzles on the three airlit graphite epoxy motors (GEMs) ■ the new 10-ft composite payload fairing ■ the new First Space Launch Squadron Operations Building (1 SLS OB) The Boeing Company will periodically update the information presented in the following pages. To this end, you are urged to promptly mail back the enclosed Readers Service Card so that you will be sure to receive any updates as they become available. Recipients are also urged to contact Boeing with comments, requests for clarification, or amplification of any information in this document. General inquiries regarding launch service availability and pricing should be directed to: Delta Launch Services Inc. Phone: (714) 896-3294 Fax: (714) 896-1186 E-mail: [email protected] Inquiries regarding the content of the Delta II Payload Planners Guide should be directed to: Delta Launch Services Customer Program Development Phone: (714) 896-5195 Fax: (714) 372-0886 E-mail: [email protected] Mailing address: Delta Launch Services c/o The Boeing Company 5301 Bolsa Avenue Huntington Beach, CA 92647-2099 U.S.A. Attn: H014-C426 VisitusatourDelta II Web site: www.boeing.com/defense-space/space/delta/delta2/delta2.htm

iii CONTENTS

INTRODUCTION 1

Section 1 LAUNCH VEHICLE DESCRIPTIONS 1-1 1.1 Delta Launch Vehicles 1-1 1.2 Launch Vehicle Description 1-2 1.3 Vehicle Axes/Attitude Definitions 1-7 1.4 Launch Vehicle Insignia 1-7

Section 2 GENERAL PERFORMANCE CAPABILITY 2-1 2.1 Launch Sites 2-1 2.2 Mission Profiles 2-1 2.3 Performance Capability 2-5 2.4 Mission Accuracy Data 2-7

Section 3 SPACECRAFT FAIRINGS 3-1 3.1 General Description 3-1 3.2 The 2.9-m (9.5 ft) Diameter Spacecraft Fairing 3-2 3.3 The 3-m (10 ft) Diameter Spacecraft Fairing 3-2 3.4 The 3-m (10L-ft) Diameter Spacecraft Fairing 3-8

Section 4 SPACECRAFT ENVIRONMENTS 4-1 4.1 Prelaunch Environments 4-1 4.2 Launch and Flight Environments 4-5

Section 5 SPACECRAFT INTERFACES 5-1 5.1 Delta Third-Stage Description 5-1 5.2 Payload Attach Fittings for Three-Stage Missions 5-1 5.3 Payload Attach Fittings for Two-Stage Missions 5-13 5.4 New PAFs 5-25 5.5 Test Fittings and Fit-Check Policy 5-37 5.6 Electrical Design Criteria 5-37

Section 6 LAUNCH OPERATIONS AT EASTERN RANGE 6-1 6.1 Organizations 6-1 6.2 Facilities 6-1 6.3 Spacecraft Transport to Launch Site 6-16 6.4 SLC-17, Pads A and B (CCAS) 6-16 6.5 Support Services 6-25 6.6 Delta II Plans and Schedules 6-28 6.7 Delta II Meetings and Reviews 6-37

v Section 7 LAUNCH OPERATIONS AT WESTERN RANGE 7-1 7.1 Organizations 7-1 7.2 Facilities 7-1 7.3 Spacecraft Transport to Launch Site 7-18 7.4 Space Launch Complex 2 7-22 7.5 Support Services 7-23 7.6 Delta II Plans and Schedules 7-34 7.7 Delta II Meetings and Reviews 7-42 Section 8 SPACECRAFT INTEGRATION 8-1 8.1 Integration Process 8-1 8.2 Documentation 8-2 8.3 Launch Operations Planning 8-3 8.4 Spacecraft Processing Requiremens 8-3 Section 9 SAFETY 9-1 9.1 Safety Requirements 9-1 9.2 Documentation Requirements 9-2 9.3 Hazardous Systems and Operations 9-3 9.4 Waivers 9-4 Appendix A DELTA MISSIONS CHRONOLOGY A-1 Appendix B NATURAL AND TRIGGERED LIGHTNING LAUNCH COMMIT CRITERIA B-1

vi FIGURES

1-1 Delta/Delta II Growth To Meet Customer Needs 1-1 1-2 Delta II Configurations 1-2 1-3 Delta 7925 Launch Vehicle 1-4 1-4 Delta 7920-10 Launch Vehicle 1-5 1-5 Delta Payload Fairings 1-8 1-6 Vehicle Axes 1-9 2-1 Typical Two-Stage Mission Profile 2-1 2-2 Typical Three-Stage Mission Profile 2-1 2-3 Typical Delta II 7320 Mission Profile—Circular Orbit Mission (Eastern Range) 2-3 2-4 Typical Delta II 7320 Mission Profile—Polar Orbit Mission (Western Range) 2-3 2-5 Typical Delta II 7925 Mission Profile—GTO Mission (Eastern Range) 2-4 2-6 Typical Delta II 7920 Mission Profile—Polar Mission (Western Range) 2-4 2-7 Delta II 7320 Vehicle, Two-Stage Perigee Velocity Capability, Eastern Range 2-8 2-8 Delta II 7320 Vehicle, Two-Stage Apogee Altitude Capability, Eastern Range 2-9 2-9 Delta II 7320 Vehicle, Two-Stage Circular Orbit Capability, Eastern Range 2-10 2-10 Delta II 7325 Vehicle, Three-Stage Perigee Velocity Capability, Eastern Range 2-11 2-11 Delta II 7325 Vehicle, Three-Stage Planetary Mission Capability, Eastern Range 2-12 2-12 Delta II 7320 Vehicle, Two-Stage Perigee Velocity Capability, Western Range 2-13 2-13 Delta II 7320 Vehicle, Two-Stage Apogee Altitude Capability, Western Range 2-14 2-14 Delta II 7320 Vehicle, Two-Stage Circular Orbit Capability, Western Range 2-15 2-15 Delta II 7320 Vehicle, Two-Stage Sun-Synchronous Orbit Capability, Western Range 2-16 2-16 Delta II 7920 Vehicle, Two-Stage Perigee Velocity Capability, Eastern Range 2-17 2-17 Delta II 7920 Vehicle, Two-Stage Apogee Altitude Capability, Eastern Range 2-18 2-18 Delta II 7920 Vehicle, Two-Stage Circular Orbit Capability, Eastern Range 2-19 2-19 Delta II 7925 Vehicle, Three-Stage Perigee Velocity Capability, Eastern Range 2-20 2-20 Delta II 7925 Vehicle, Three-Stage Apogee Altitude Capability, Eastern Range 2-21 2-21 Delta II 7925 Vehicle, Three-Stage GTO Inclination Capability, Eastern Range 2-22

vii 2-22 Delta II 7925 Vehicle, Three-Stage Planetary Mission Capability Eastern Range 2-23 2-23 Delta II 7920 Vehicle, Two-Stage Perigee Velocity Capability, Western Range 2-24 2-24 Delta II 7920 Vehicle, Two-Stage Apogee Altitude Capability, Western Range 2-25 2-25 Delta II 7920 Vehicle, Two-Stage Circular Orbit Capability, Western Range 2-26 2-26 Delta II 7920 Vehicle, Two-Stage Sun-Synchronous Orbit Capability, Western Range 2-27 2-27 Delta II 7925 Vehicle, Three-Stage Perigee Velocity Capability, Western Range 2-28 2-28 Delta II 7925 Vehicle, Three-Stage Apogee Altitude Capability, Western Range 2-29 2-29 Delta II 7925 Vehicle, Three-Stage GTO Deviations, Eastern Range 2-30 3-1 Delta 2.9-m (9.5 ft) Payload Fairing 3-3 3-2 Profile, 2.9-m (9.5 ft) Diameter Fairing 3-4 3-3 Spacecraft Envelope, 2.9-m (9.5 ft) Diameter Fairing, Three-Stage Configuration (3712 PAF) 3-5 3-4 Spacecraft Envelope, 2.9-m (9.5 ft) Diameter, Two-Stage Configuration (6915 PAF) 3-6 3-5 Profile, 3-m (10 ft) Composite Fairing 3-7 3-6 Spacecraft Envelope, 3-m (10 ft) Diameter Fairing, Three-Stage Configuration (3712 PAF) 3-9 3-7 Spacecraft Envelope, 3-m (10-ft) Diameter Fairing, Two-Stage Configuration (6915 PAF) 3-10 3-8 Profile, 3m (10-ft) 10L Composite Fairing 3-11 3-9 Spacecraft Envelope, 3-m (10-ft) Diameter 10L Fairing, Three-Stage Configuration (3712) PAF 3-12 3-10 Spacecraft Envelope, 3-m (10-ft) Diameter 10L Fairing, Two-Stage Configuration (6915) PAF 3-13 4-1 Payload Air Distribution System 4-1 4-2 Delta II Payload Fairing Compartment Absolute Pressure Envelope 4-6 4-3 Predicted Maximum Internal Wall Temperatures and Internal Surface Emittances (9.5-ft Fairing) 4-7 4-4 Predicted Maximum Internal Wall Temperatures and Internal Surface Emittances (10-ft Fairing) 4-8 4-5 Predicted Star 48B Plume Radiation at the Spacecraft Separation Plane Versus Radial Distance 4-9 4-6 Predicted Star 48B Plume Radiation at the Spacecraft Separation Plane Versus Burn Time 4-10

viii 4-7 Predicted Star 48B Motor Case Soakback Temperature 4-10 4-8 Axial Steady-State Acceleration at MECO Versus Second-Stage Payload Weight 4-11 4-9 Axial Steady-State Acceleration at Third-Stage Burnout 4-12 4-10 Predicted Delta II 7920 and 7925 9.5-ft Fairing Spacecraft Acoustic Environment 4-13 4-11 Predicted Delta II 7920 and 7925 10-ft Fairing Spacecraft Acoustic Environment 4-14 4-12 Maximum Flight Spacecraft Interface Shock Environment—3712A, 3712B, 3712C Payload Attach Fitting 4-16 4-13 Maximum Flight Spacecraft Interface Shock Environment—6306 Payload Attach Fitting 4-16 4-14 Maximum Flight Spacecraft Interface Shock Environment—6019 and 6915 Payload Attach Fitting 4-17 4-15 Maximum Flight Spacecraft Interface Shock Environment—5624 Payload Attach Fitting 4-17 4-16 Delta II Spin Rate Capability 4-23 4-17 Maximum Expected Angular Acceleration Versus Spin Rate 4-23 5-1 Delta II Upper Stage 5-1 5-2 3712 PAF Assembly 5-3 5-3 Capability of the 3712 PAF Configuration 7925 Vehicle 5-3 5-4 3712 PAF Detailed Assembly 5-4 5-5 3712A PAF Detailed Dimensions 5-5 5-6 3712A PAF Spacecraft Interface Dimensional Constraints 5-5 5-7 3712A PAF Spacecraft Interface Dimensional Constraints (Views C, D, E, and Section B-B) 5-6 5-8 3712B PAF Detailed Dimensions 5-7 5-9 3712B PAF Spacecraft Interface Dimensional Constraints 5-7 5-10 3712B PAF Spacecraft Interface Dimensional Constraints (Views C, D, E, and Section B-B) 5-8 5-11 3712C PAF Detailed Dimensions 5-9 5-12 3712C PAF Spacecraft Interface Dimensional Constraints 5-9 5-13 3712C PAF Spacecraft Interface Dimensional Constraints (Views B-B, C, D, and E)5-10 5-14 3712 PAF Interface 5-11 5-15 3712 Clamp Assembly and Spring Actuator 5-12 5-16 3712 PAF Bolt Cutter Detailed Assembly 5-13 5-17 6019 PAF Assembly 5-14

ix 5-18 Capability of the 6019 PAF on 7920 Vehicle 5-15 5-19 6019 PAF Spacecraft Interface Dimensional Constraints 5-16 5-20 6019 PAF Detailed Assembly 5-17 5-21 6019 PAF Spacecraft Assembly 5-18 5-22 6019 PAF Detailed Dimensions 5-18 5-23 6915 PAF 5-19 5-24 Capability of the 6915 PAF on 7920 Vehicle 5-19 5-25 6915 PAF Detailed Assembly 5-21 5-26 6915 PAF Spacecraft Interface Dimensional Constraints 5-22 5-27 6915 PAF Spacecraft Assembly 5-23 5-28 6915 PAF Detailed Dimensions 5-23 5-29 Actuator Assembly Installation–6915 PAF 5-24 5-30 6306 PAF Assembly 5-25 5-31 6306 PAF Detailed Dimensions 5-26 5-32 6306 PAF Detailed Dimensions 5-27 5-33 6306 PAF Spacecraft Interface Dimensional Constraints 5-28 5-34 6306 PAF Spacecraft Interface Dimensional Constraints 5-29 5-35 Capability of the 6306 PAF on 7920 Vehicle 5-30 5-36 6306 Secondary Latch 5-31 5-37 5624 PAF Detailed Assembly 5-32 5-38 5624 PAF Detailed Dimensions 5-33 5-39 5624 PAF Clamp Assembly and Spring Actuator 5-34 5-40 5624 PAF Spacecraft Interface Dimensional Constraints 5-35 5-41 5624 PAF Spacecraft Interface Dimensional Constraints 5-36 5-42 Capability of 5624 PAF on 7900 Vehicle 5-37 5-43 Typical Delta II Wiring Configuration 5-39 5-44 Typical Payload-to-Blockhouse Wiring Diagram for Three-Stage Missions at SLC-17 5-41 5-45 Typical Payload-to-Blockhouse Wiring Diagram for Three-Stage Missions at SLC-2 5-42 5-46 Typical Spacecraft Umbilical Connector 5-43 5-47 Spacecraft/Fairing Umbilical Clearance Envelope 5-43 5-48 Typical Spacecraft Separation Switch and PAF Switch Pad 5-44 5-49 Spacecraft/Pad Safety Console Interface for SLC-17 5-44 5-50 Spacecraft/Pad Safety Console Interface for SLC-17-Operations Building Configuration 5-45

x 6-1 Organizational Interfaces for Commercial Users 6-2 6-2 Astrotech Site Location 6-3 6-3 Astrotech Complex Location 6-3 6-4 Astrotech Building Locations 6-3 6-5 First Level Floor Plan, Building 1/1A, Astrotech 6-5 6-6 Second Level Floor Plan, Building 1/1A, Astrotech 6-6 6-7 Building 2 Detailed Floor Plan, Astrotech 6-8 6-8 Building 3 Detailed Floor Plan, Astrotech 6-9 6-9 Building 4 Detailed Floor Plan, Astrotech 6-9 6-10 Building 5 Detailed Floor Plan, Astrotech 6-10 6-11 Building 6 Detailed Floor Plan, Astrotech 6-10 6-12 Delta Spacecraft Checkout Facilities 6-11 6-13 Cape Industrial Area 6-11 6-14 Building AE Floor Plan 6-12 6-15 Building AE Mission Director Center 6-13 6-16 SAEF-2 Floor Plan 6-14 6-17 Electrical-Mechanical Testing Facility Floor Plan 6-16 6-18 Delta II Upper Stage Assembly Ground-Handling Can and Transporter 6-17 6-19 SLC-17, CCAS 6-19 6-20 SLC-17—Aerial View, Facing South 6-19 6-21 Environmental Enclosure Work Levels 6-20 6-22 Level 9A Floor Plan, Pads 17A and 17B 6-21 6-23 Level 9B Floor Plan, Pads 17A and 17B 6-21 6-24 Level 9C Floor Plan, Pads 17A and 17B 6-22 6-25 Blockhouse Console Locations 6-22 6-26 Spacecraft-to-Blockhouse Junction Box 6-23 6-27 Spacecraft Customer Accomodations—Launch Control Center 6-23 6-28 Interface Overview—Spacecraft Control Rack in 1 SLS OB 6-24 6-29 Launch Decision Flow for Commercial Missions—Eastern Range 6-25 6-30 Delta II 792X Ground Wind Velocity Criteria, SLC-17 6-26 6-31 Typical Mission Plan 6-29 6-32 Typical Spacecraft Weighing (F-11 Day) 6-29 6-33 Typical Mating of Spacecraft and Third Stage (F-10 Day) 6-30 6-34 Typical Final Spacecraft Third-Stage Preparations (F-9 Day) 6-30 6-35 Typical Installation of Transportation Can (F-8 Day) 6-31 6-36 Typical Spacecraft Erection (F-7 Day) 6-31

xi 6-37 Typical Flight Program Verification and Stray Voltage Checks (F-6 Day) 6-32 6-38 Typical Ordnance Installation and Hookup (F-5 Day) 6-32 6-39 Typical Fairing Installation (F-4 Day) 6-33 6-40 Typical Fairing Installation (F-4 Day) 6-33 6-41 Typical Propellant Loading Preparations (F-3 Day) 6-34 6-42 Typical Second-Stage Propellant Loading (F-2 Day) 6-34 6-43 Typical Beacon, Range Safety, and Class A Ordnance (F-1 Day) 6-35 6-44 Typical Delta Countdown (F-0 Day) 6-35 6-45 Typical Terminal Countdown (F-0 Day) 6-36 7-1 Launch Base Organization at VAFB for Commercial Launches 7-2 7-2 Vandenberg Air Force Base 7-3 7-3 Spacecraft Support Area 7-4 7-4 Spacecraft Laboratory (Building 836) 7-5 7-5 Launch Vehicle Data Center, Building 836, Vandenberg AFB 7-6 7-6 NASA Building 840 7-7 7-7 Mission Director Center—Building 840 7-8 7-8 Hazardous Processing Facility 7-9 7-9 Spin Test Building (Building 1610) 7-10 7-10 Control Room Building 1605 Floor Plan 7-11 7-11 Astrotech Payload Processing Area 7-12 7-12 Astrotech Payload Processing Facility—Building 1032 7-13 7-13 California Spaceport—Plan View of the IPF 7-14 7-14 California Spaceport—IPF Cross-Sectional View 7-14 7-15 California Spaceport—Cutaway View of the IPF (looking South) 7-15 7-16 California Spaceport—Processing Area 7-19 7-17 California Spaceport—Level 89 Technical Support Area 7-20 7-18 California Spaceport—Level 101 Technical Support Area 7-20 7-19 Upper-Stage Assembly Ground Handling Can and Transporter 7-21 7-20 Space Launch Complex-2, Vandenberg Test Center 7-22 7-21 Space Launch Complex SLC-2, VAFB—Aerial View Looking East 7-23 7-22 SLC-2 MST/FUT Elevation 7-24 7-23 Level 5 of SLC-2 Mobile Service Tower—Plan View 7-25 7-24 Level 6 of SLC-2 Mobile Service Tower—Plan View 7-26 7-25 Spacecraft Work Levels in SLC-2 Mobile Service Tower—VAFB 7-27 7-26 Whiteroom Elevations and Hook Heights—SLC-2 Mobile Service Tower 7-28 7-27 SLC-2 Blockhouse 7-29

xii 7-28 SLC-2 Blockhouse Control Room 7-29 7-29 Spacecraft Blockhouse Console—Western Range 7-30 7-30 Launch Decision Flow for Commercial Missions—Western Range 7-31 7-31 Delta II 792X Ground Wind Velocity Criteria, SLC-2 7-32 7-32 Typical Mission Plan 7-35 7-33 Typical Spacecraft Weighing (F-12 Day) 7-35 7-34 Typical Spacecraft/PAM Mate (F-11 Day) 7-36 7-35 Typical Spacecraft/PAM Final Preparations (F-10 Day) 7-36 7-36 Typical Transportation Can Installation (F-9 Day) 7-37 7-37 Typical Spacecraft Erection (F-8 Day) 7-37 7-38 Typical Flight Program Verification and Saftety Stray Voltage Checks (F-7 Day) 7-38 7-39 Typical Ordnance Installation (F-6 Day) 7-38 7-40 Typical Fairing Installation (F-5 Day) 7-39 7-41 Typical Fairing Finaling (F-4 Day) 7-39 7-42 Typical Fairing Finaling (F-3 Day) 7-40 7-43 Typical Second-Stage Propellant Loading (F-2 Day) 7-40 7-44 Typical Beacon, Range Safety, and Class A Ordnance (F-1 Day) 7-41 7-45 Typical Delta Countdown (F-0 Day) 7-41 8-1 Mission Integration Process 8-1 8-2 Typical Delta II Agency Interfaces 8-2 8-3 Typical Document Interfaces 8-3 8-4 Typical Integration Planning Schedule 8-14 8-5 Launch Operational Configuration Development 8-15 9-1 General Safety Documentation Flow for Commercial Missions 9-3

xiii TABLES

1-1 Delta Four-Digit Designation 1-3

1-2 Delta II Vehicle Description 1-6

1-3 Delta II Vehicle Characteristics 1-6

2-1 Delta II Typical Eastern Range Event Times 2-5

2-2 Delta II Typical Western Range Event Times 2-5

2-3 Mission Capabilities 2-6

3-1 Typical Acoustic Blanket Configurations 3-1

4-1 Eastern Range Facility Environments 4-2

4-2 Western Range Facility Environments 4-2

4-3 Delta II Transmitter Characteristics 4-3

4-4 Cleanliness Level Definition 4-4

4-5 Spacecraft CG Limit-Load Factors (g) 4-13

4-6 Spacecraft Acoustic Environment Figure Reference 4-13

4-7 Sinusoidal Vibration Levels 4-14

4-8 Spacecraft Interface Shock Environment Figure Reference 4-15

4-9 Acoustic Test Levels, Delta II 7925 9.5-ft Fairing, Three-Stage

Mission, 3.0-in. Blanket Configuration 4-19

4-10 Acoustic Test Levels, Delta II 7920 9.5-ft Fairing, Two-Stage

Mission, 3.0-in. Blanket Configuration 4-19

4-11 Acoustic Test Levels, Delta II 7920 and 7925 10-ft Fairing, Two- and Three-Stage Mission, 3.0-in. Blanket Configuration 4-20

4-12 Sinusoidal Vibration Acceptance Test Levels 4-20

4-13 Sinusoidal Vibration Qualification Test Levels 4-20

4-14 Sinusoidal Vibration Protoflight Test Levels 4-20

4-15 Third Stage Mass Properties 4-24

4-16 NCS Nominal Characteristics 4-24

xv 5-1 Notes Used in Configuration Drawings 5-2

5-2 Maximum Clamp Assembly Preload 5-2

5-3 Separation Clamp Assemblies 5-37

5-4 One-Way Line Resistance 5-40

5-5 Disconnect Pull Forces (Lanyard Plugs) 5-42

5-6 Disconnect Forces (Rack and Panel Connectors) 5-42

5-7 Disconnect Forces (Bayonet-Mate Lanyards) 5-43

6-1 Test Console Items 6-16

7-1 Vehicle Checkout Facility 7-16

7-2 Vehicle Checkout Area 7-16

7-3 Payload Checkout Cells Capabilities 7-17

7-4 Transfer Tower Area 7-18

7-5 Fairing Storage and Assembly Area 7-18

7-6 Payload Processing Room 6902 7-19

7-7 Payload Processing Room (PPR) 8910 1-19

8-1 Space Agency Data Requirements 8-4

8-2 MDA Program Documents 8-4

8-3 Required Documents 8-5

8-4 Spacecraft Questionnaire 8-9

8-5 Typical Spacecraft Launch Site Test Plan 8-12

8-6 Data Required for Orbit Parameter Statement 8-13

8-7 Spacecraft Checklist 8-16

9-1 Safety Document Applicability 9-2

xvi EED electro-explosive device GLOSSARY EMF electromagnetic field 1 SLS OB First Space Launch Squadron EMI electromagnetic interference Operations Building EMT Electrical-Mechanical Testing Facility 30 SW 30th Space Wing ER Eastern Range 45 SW 45th Space Wing ESA Explosive Safe Area A/C air conditioning E/W east/west ACS attitude control system EWR Eastern Western Regulation ADOTS Advanced Delta ordnance test set FAA Federal Aviation Administration AFB Air Force Base FO fiber-optic AGE aerospace ground equipment FRR flight readiness review AKM apogee kick motor FS first stage AL air-lit FSAA fairing storage and assembly area ALC advanced launch control system FUT fixed umbilical tower AMP Ampers GC guidance computer ANSI American National Standard Institute GC&NS guidance, control, and ARAR Accident Risk Assessment Report navigation system ASO Astrotech Space Operations GEM graphite-epoxy motor ATP authority to proceed GL ground-lit AUV avionics upgraded vehicle GMT Greenwich meantime AWG American Wire Gauge GN2 gaseous nitrogen BAS breathing air supply GSE ground support equipment B/H blockhouse GSFC Goddard Space Flight Center CAD computer-aided design GTO geosynchronous transfer orbit CCAS Cape Canaveral Air Station H2 hydrogen CG center of gravity HDBK handbook C/O checkout H/H hook height CRD command receiver decoder HPF hazardous processing facility CWA clean work area HPTF high pressure test facility DBL Dynamic Balance Laboratory I/F interface DID data item descriptions IIP instantaneous impact point DIGS Delta Inertial Guidance System IPF integrated processing facility DMCO Delta mission checkout J-box junction box DOT Department of Transportation KMI KSC Management Instruction DRIMS Delta redundant inertial KSC Kennedy Space Center measurement system LEO low Earth orbit DTO detailed test objectives LCC Launch Control Center E&O engineering and operations LCE launch control equipment

EAL Entry Authority List LO2 liquid oxygen

xvii LOCC launch operations control center PCS probability of command shutdown LOP Launch Operations Plan PEA payload encapsulation area LOX Liquid Oxygen PGOC payload ground operations contract LPD Launch Processing Document PHE propellant handler's ensemble LRR launch readiness review PI program introduction LSRR launch site readiness review PLF payload fairing LSSM Launch Site Support Manager PMA Preliminary Mission Analysis LSTP Launch Site Test Plan P/N part number LV launch vehicle PPF payload processing facility LVDC Launch Vehicle Data Center PPG Payload Planners Guide MD Mission Director PPR payload processing room MDA McDonnell Douglas Aerospace MDC Mission Director Center PPRD Payload Processing Requirements MDC McDonnell Douglas Corporation Document MECO main-engine cutoff PRD Program Requirements Document MIC meets intent certification PSM Program Support Manager MMS multimission modular spacecraft PSP Program Support Plan MOI moments of inertia PSSC pad safety supervisor's console MRTB Missile Research Test Building PWU portable weigh unit MSPSP Missile Systems Prelaunch QD quick disconnect Safety Package RACS redundant attitude control system MSR Mission Support Request RCS reaction control system RF radio frequency MST mobile service tower RFA radio frequency application NASA National Aeronautics and Space RFI radio frequency interference Administration RGA rate gyro assembly NCS nutation control system RIFCA Eedundant Inertial Flight Control Assembly NDTL Nondestructive Testing Laboratory N/S north/south ROS Range Operations Specialist NOAA National Oceanographic and RS range safety Astronautic Agency S&A safe and arm OASPL overall sound pressure level SAB Sterilization and Assembly Building OLS Orbital Launch Services SAEF 2 Spacecraft Assembly and Encapsulation OR Operations Requirement Facility Number 2 OVS operational voice system SECO second-stage engine cutoff PAM payload assist module SLC Space Launch Complex P&C power and control S/C spacecraft PAF payload attach fitting SLS Space Launch Squadron PCC payload checkout cell S/M Solid Motor PCM pulse coded modulated SMC Space and Missile Center xviii SOB squadron operations building UV ultraviolet SOP standard operating procedure VAB Vertical Assembly Building SR&QA safety, reliability, and quality VAC volts alternating current assurance VAFB Vandenberg Air Force Base SRM solid rocket motor VC visible cleanliness SS Second Stage VCA vehicle checkout area STD standard VCF vehicle checkout facility STG stage VCR1 vehicle control rack 1 STS Space Transportation System VCR2 vehicle control rack 2 SW Space Wing VDC volts direct current TIM technical interchange meeting VEH vehicle TM telemetry VIM Vehicle Information Memorandum TMR telemetry control rack VLD voice direct line TMS telemtry system VM video monitor TWX telex VOS vehicle on stand Typ typical W/D Walkdown UDS Universal Document System W/O without UMB umbilical WR Western Range USAF United States Air Force

xix INTRODUCTION successes of the Delta launch vehicle stem from its evolutionary design, which can be upgraded to This Planners Guide is provided by the McDonnell meet the needs of the user community while main- Douglas Corporation (MDC) to familiarize poten- taining the highest reliability of any Western launch tial customers with the Delta II launch services. The vehicle. guide describes the Delta II, its background and The Delta provides flexibility by maintaining heritage, and its performance capabilities. Space- launch pads at both the eastern and western ranges. craft interface constraints and the environments Two launch pads at Cape Canaveral Air Station that the spacecraft will experience during launch (CCAS) in Florida have been in service longer than are defined. Facilities, operations, and operational the pads used by any other launch system and have constraints are described, as well as the documenta- been regularly upgraded to meet the increasingly tion, integration, and procedural requirements that rigorous spacecraft support requirements of our are associated with preparing for and conducting a customers. The pads are open to both commercial launch. and government customers without restriction. The Delta II is the newest, most powerful version Maintenance, mission modifications, and launch of the highly regarded Delta series of launch ve- preparations may be conducted at one pad without hicles originally developed by the National Aero- impacting operations at the other. This arrange- nautics and Space Administration (NASA). In over ment allows MDC to provide schedule and launch- three decades of use, Delta has achieved an average period flexibility and to minimize schedule risk. of one launch every 60 days. The long life and many Having an additional pad at Vandenberg Air Force

DAC123862 H31195

Delta II early morning final launch preparations—Eastern Range 1 DAC113761 H31194

Delta II launch—liftoff at Eastern Range

2 Base (VAFB) in California provides mission flex- DAC123861 H31202 ibility. The Delta II is launched from one of the dedicated pads at Space Launch Complex (SLC) 17 at CCAS, for missions requiring low- and medium- inclination orbits; and from SLC-2, a dedicated Delta launch pad at VAFB, for high-inclination orbits. Our demonstrated high launch rate is made possible by streamlined procedures, experienced personnel, and excellent working relationships with all Range regulatory and safety agencies. The Delta launch team is one of the most experienced in the world. It has a reputation for responding quickly to mission-specific requirements including providing services for the highly specialized payloads of NASA, USAF, and commercial agencies. MDC approaches each mission by fully involving its customers in the integration and launch preparations process. This team approach is aimed at producing quality results and mission success. As a commercial launch service provider, MDC acts as the sole agent for the user in interfacing with the United States Air Force (USAF), National Aeronautics and Space Administration (NASA), Department of Transportation (DOT), and Astrotech Space Operations Company, as well as any other agency necessary when other commercial or government facilities are engaged for spacecraft processing. A commercialization agreement Delta II contrail—seconds after liftoff with the USAF and NASA provides MDC use of the launch facilities and services in support of Delta II conditions. At the appropriate time, a mission commercial vehicles. manager will be assigned from the Delta Program Specific Delta II personnel will be assigned to the Office to coordinate all matters related to the inte- user’s program at McDonnell Douglas Aerospace gration of the spacecraft with the Delta II launch (MDA) in Huntington Beach, California. In the vehicle and facilities; interface with government or early, precontractual stages, a marketing manager other agencies; and documentation, planning, and will coordinate matters related to launch capability, scheduling requirements. spacecraft interfaces, integration, and other engi- The Delta team addresses each customer’s con- neering issues, as well as contractual terms and cerns and requirements individually, employing a 3 meticulous, systematic, user-specific process that Delta II operations from coast to coast provide covers advance mission planning and analysis of proven efficiency in production, integration, and spacecraft design; coordination of systems inter- launch services. More than 80% of the Delta II face between spacecraft and Delta II; processing of fabrication and subassembly occurs in Huntington all necessary documentation, including government Beach, California. The final assembly takes place requirements; prelaunch systems integration and in Pueblo, Colorado, and vehicle checkout is con- checkout; launch site operations dedicated exclu- ducted at either CCAS or VAFB. sively to the user’s schedule and needs; and post- The Delta II offers a dedicated launch service, the flight analysis. benefit of a launch team committed to the users’ The Delta team works closely with its customers payload, and a mission profile and launch window to define optimum performance for the mission designed specifically for a single payload. This payload(s). In many cases we can provide innova- dedicated manifest results in better cost control and tive performance trades to augment the perfor- greater overall efficiency, plus the advantages of a mance shown in Section 2. Our Delta team also has simplified integration process, on-time launch as- extensive experience in supporting customers around surance, and straightforward flight operations and the globe. This demonstrated capability to utilize control. Coupled with this are the McDonnell the flexibility of the Delta vehicle and design team, Douglas commitment to excellence and proven together with our experience in supporting world- dependability, which have given our customers the wide customers, makes Delta the ideal choice as a highest assurance of a successful launch campaign launch service provider. for more than three decades.

DAC61666 H31193

Eastern Range Launch Complex 17 4 Md Section 1 vehicle improvement to meet customer needs led to LAUNCH VEHICLE DESCRIPTIONS the Delta II vehicle, which now provides a capability of over 1869 kg (4120 lb) to GTO (Figure 1-1). This section provides an overall description of The Delta has compiled an impressive record of the Delta II launch vehicle and its major compo- successful launches for more than three decades. nents. In addition, the Delta vehicle designations During this period it has demonstrated exceptional are explained. reliability in the launching of satellites for commu- 1.1 DELTA LAUNCH VEHICLES nication and navigation, meteorology, science, and The Delta launch vehicle program was initiated Earth observation for government and commercial in the late 1950s by the National Aeronautics and users worldwide. The launch history presented in Space Administration (NASA) with McDonnell the Appendix (Delta Missions Chronology) sum- Douglas as the prime contractor. McDonnell Dou- marizes the Delta tradition of success. It can be seen glas developed an interim space launch vehicle that in the 18 years prior to the publication of this using a modified Thor as the first stage and Van- document, the Delta has had a launch success rate of guard components as the second and third stages. over 98%. This record is the best of any launch The vehicle was capable of delivering a payload of system currently available and is a strong argument 54 kg (120 lb) to geosynchronous transfer orbit for choosing the Delta II as the launcher for your (GTO). The McDonnell Douglas commitment to mission.

H54532.3 Avionics Upgrades, 10-ft-dia Fairing, Ordnance Thrusters, Extended Air-Lit GEMs Nozzles RS-27A Main Engine, Graphite/Epoxy SRMs 9.5-ft-dia Payload Fairing, 12-ft Stretch for Propellant Tank, Castor IVA SRMs 1996 New 2nd Stage (4400) Payload Assist Module 3rd Stage Delta II Delta II 7925 7925 1814 Delta Redundant Inertial Measuring System (4000) Engine Servo-System Electronics Package 1633 Castor IV SRMs (3600) RS-27 Main Engine, 8-ft Payload Fairing, Isogrid Main System 9 Castor SRMs, Delta Inertial Guidance System Delta II 1451 6925 (3200) 6 Castor SRMs 3920/ Stretched Propellant Tank 1270 PAM-D Upgraded 3rd Stage (2800) 3910/ 3 Castor II SRMs 1089 PAM-D 5-ft-dia Payload Fairing (2400) Revised MB-3 Main Engine 907 3 Castor I SRMs 3914 (2000) Revised MB-3

Payload to GTO in kilograms (lb) kilograms in GTO to Payload 726 Main Engine and (1600) 3rd Stage 2914 MM6 904 544 C D EJ (1200) Delta

363 (800)

181 (400)

0 1960 1963 1964 1965 1968 1969 19701971 1973 1975 1980 1982 1989 1990 199 5 Figure 1-1. Delta/Delta II Growth To Meet Customer Needs 1-1 The Delta II offers both two- and three-stage and the Alliant lightweight, GEM solid rocket strap- vehicles; a choice of 2.9- and 3-m (9.5- and 10-ft) ons. The booster is available with either nine (792X diameter fairings; and three, four, or nine solid designation), four (742X designation), or three (732X motor configurations to meet the needs of our designation) GEM strap-ons. The three-stage 7925 customers (Figure 1-2). and the two-stage 7920-10 vehicles shown in The four-digit system used to identify Delta Figures 1-3 and 1-4, respectively, are representative configurations is explained in Table 1-1. of the series. Delta II characteristics are summarized in Tables 1-2 and 1-3. 1.2 LAUNCH VEHICLE DESCRIPTION The major elements of the Delta II launch vehicle 1.2.1 First Stage are the first stage and its thrust augmentation solid The first-stage subassemblies include the motors, the second stage, the third stage and spin interstage, fuel tank, centerbody, liquid oxygen (LO2) table, and the payload fairing (PLF). The Delta tank, and engine section. 7925 has a nominal overall length of 38.2 m (125.2 The Rocketdyne RS-27A main engine has a 12:1 ft) and a diameter of 2.4 m (8 ft) for the core vehicle. expansion ratio and employs a turbine/turbopump The vehicle is an integrated system with a proven and a regeneratively cooled thrust chamber and heritage of reliability, operability, and producibility. nozzle. The thrust chamber and nozzle are hydrau- The current 7000 series booster configuration lically gimbaled to provide pitch and yaw control. uses an RS-27A engine with a 12:1 expansion ratio Two Rocketdyne vernier engines provide roll con-

03296REU7 2.9 3.0 3.0 3.0 9.5 10.0 10.0 10.0 m ft 8.5 8.9 9.3 8.9 27.8 29.1 30.4 29.1 Paste

38.2 38.6 38.9 38.6 125.2 126.5 127.8 126.5

Graphite Epoxy Motors (GEMs)

12:1 Main Engine Delta 7925 Delta 7925-10 Delta 7925-10L Delta 7320-10 Expansion Ratio Delta 7420-10

Figure 1-2. Delta II Configurations 1-2 Table 1-1. Delta Four-Digit Designation LO2 tanks houses the first-stage electronic compo- Digit Indicates Examples 1st Type of 7 Ð GEM solid motor nents on hinged panels for easy checkout access and first-stage augmentation; extra-long enhanced maintainability. engine and extended tank; modified solid rocket RS-27A engine; 12:1 The interstage, located between the first stage motors nozzle ratio 2nd Number of 9 Ð Nine solid rocket motors and second-stage miniskirt, carries the loads from solid rocket 3 Ð Three solid rocket motors the second stage and fairing to the first stage. The motors 4 Ð Four solid rocket motors 3rd Type of 2 Ð Aerojet AJ10-118K interstage houses the second stage and contains second engine stage range safety antennas, an exhaust vent for the fair- 4th Type of 0 Ð No third stage ing cavity, and six guided spring actuators to sepa- third stage 5 Ð STAR 48B third stage (4,430 lb propellant rate the second stage from the first stage. maximum) Dash Type of None Ð Standard 9.5-ft-diameter, no. fairing 27.8-ft-long fairing 1.2.2 Second Stage -10 Ð 10-ft diameter, 29.1-ft- long-fairing The second stage is powered by the proven Aerojet -10L Ð 10-ft diameter, 30.4-ft long fairing (under AJ10-118K engine and includes fuel and oxidizer development) tanks that are separated by a common bulkhead. Example: Delta 7925-10 Digit Indicates The simple, reliable start and restart operation re- 7 GEM solid motor augmentation; extra-long extended tank; modified RS-27A engine; 12:1 quires only the actuation of a bipropellant valve to nozzle ratio release the pressure-fed hypergolic propellants with 9 Nine solid rocket motors 2 Aerojet AJ10-118K engine no need for a turbopump or an ignition system. 5 STAR 48B third stage (4,430 lb propellant maximum) Typical Delta two- and three-stage missions utilize -10 10-ft diameter, 29.1-ft-long fairing two second-stage starts, but the restart capability has been used as many as six times on a single mission. During powered flight, the second-stage trol during main-engine burn and attitude control hydraulic system gimbals the engine for pitch and between main-engine cutoff (MECO) and second- yaw control. A redundant attitude control system stage separation. (RACS) using nitrogen gas provides roll control. The standard vehicle configuration includes nine The RACS also provides pitch, yaw,and roll control Alliant graphite-epoxy motors (GEMs) to augment during unpowered flight. The guidance system is the first-stage performance. Six of these motors are installed in the forward section of the second stage. ignited at liftoff, and the remaining three motors have extended nozzles and are ignited in flight after 1.2.3 Third Stage burnout of the first six. Ordnance for the motor The third stage consists of a STAR 48B solid ignition and separation systems is completely re- rocket motor (SRM), a payload attach fitting (PAF) dundant. The 732X and 742X vehicles include with a nutation control system (NCS), and a spin either three or four GEMs, all of which are ignited table containing small rockets for spin-up of the at liftoff. third stage and spacecraft. This stack mates to the

The LO2 tank, fuel tank, and interstage are con- top of the Delta second stage. structed of aluminum isogrid shells and aluminum The flight-proven STAR 48B SRM is produced tank domes. The centerbody between the fuel and by the Thiokol Corporation. The motor was devel- 1-3 H54539.2

Spacecraft Fairing Fairing

Payload Attach Fitting

Third-Stage Motor

Third-Stage Motor Separation Clamp Bands Spin Table

Guidance Fairing Electronics Access Door Second-Stage Miniskirt and Support Truss

Second Stage Helium Spheres (3) Nitrogen Sphere

Interstage

Fuel Tank Centerbody Section

Thrust Augmentation Solids First-Stage Oxidizer Tank

Figure 1-3. Delta 7925 Launch Vehicle 1-4 H54535.2

Spacecraft

Fairing

Payload Attach Second-Stage Fitting Miniskirt and Support Truss Guidance Electronics

Fairing

Second Stage

Helium Spheres (3) Nitrogen Sphere

Interstage

First Stage Fuel Tank Centerbody Section

Thrust Augmentation Solids First Stage Oxidizer Tank

Figure 1-4. Delta 7920-10 Launch Vehicle 1-5 Table 1-2. Delta II Vehicle Description Vehicle Configuration ■ Ascent environment 7320 7925 ● Liftoff mass 151,740 kg (334,525 lb)1 231,870 kg (511,190 lb)2 ● Liftoff thrust 1,970 kN (443,250 lb) 3,110 kN (699,250 lb) ● Maximum dynamic pressure 62,720 N/m2 (1310 psf) 62,720 N/m2 (1310 psf) ● Maximum steady-state acceleration3 6.65 g 6.25 g 1. Based on spacecraft mass of 2868 kg (6324 lb) 2. Based on spacecraft mass of 1869 kg (4120 lb) 3. Occurs at MECO for both example vehicle configurations, three-sigma high value quoted M029, T037, 12/8/95, 02:14 PM oped from a family of high-performance apogee motor, and to separate the spacecraft following and perigee kick motors made by Thiokol. motor burn.

Our flight-proven NCS maintains orientation of 1.2.4 Payload Attach Fittings the spin-axis of the SRM/spacecraft during third- The spacecraft mates to the Delta using an MDA- stage flight until just prior to spacecraft separation. provided payload attach fitting (PAF). A variety of The NCS uses monopropellant hydrazine that is PAFs are available for two- and three-stage mis- prepressurized with helium. This simple system sions to meet payload needs. The spacecraft separation systems are incorporated into the launch ve- has inherent reliability with only one functioning hicle PAF and include clampband separation sys- component and leak-free design. tems or attach bolt systems as required. The PAFs An ordnance sequence system is used to release and separation systems are discussed in greater the third stage after spin-up, to fire the STAR-48B detail in Section 5.

Table 1-3. Delta II Vehicle Characteristics Strap-on solids First stage Second stage Third stage

Length (m/ft) 13.0/42.5 26.1/85.6 6.0/19.6 2.0/6.7 Diameter (m/ft) 1.0/3.3 2.4/8 2.4/8 1.2/4.1 Total weight (kg/lb) 13,082/28,840 (GL)* 101,796/224,420 6,954/15,331 2,217/4,887 13,204/29,110 (AL)** Engine/motor GEM RS-27A AJ10-118K STAR-48B Manufacturer Alliant Rocketdyne Aerojet Thiokol Quantity 9, 3, or 4 1 1 1 or 0 Propellants Solid LO2/RP-1 N2O4/A-50 Solid Propellant weight (kg/lb) 11,765/25,937 each 96,118/211,902 6,004/13,236 2,009/4,430 Thrust (N/lb)† Ð SL 446,023/100,270 each 889,644/200,000 — — Ð VAC 499,180/112,220 each (GL)* 1,085,811/244,088 43,657/9,815 66,367/14,920 516,216/116,050 each (AL)**

Isp (sec)† Ð SL 245.4 254.2 — — Ð VAC 274.0 (GL)* 301.7 319.2 292.2 283.4 (AL)** Burn time (sec) 63.3 260.5 431.6 87.1 Propellant temperature 22.8/73 28.9/84 15.6/60 15.6/60 (°C/°F) Expansion ratio 10 12 65 54.8 Ê*Ground lit **Air lit (extended nozzle) †Average during the burn 1-6 M029, T002, 5 /10/96, 10:07 AM 1.2.5 Payload Fairings an electronics package (E-pack) that interfaces with The Delta launch vehicle offers the user a 2.9-m RIFCA through the P&C box to control the vehicle (9.5-ft) diameter skin-and-stringer center section attitude, and a PCM telemetry system (T/M) that fairing (bisector) as well as a 3-m (10-ft) diameter provides vehicle system performance data. (bisector) composite version. Each of these fairings The RIFCA contains the basic control logic that (Figure 1-5) can be used on either two-stage or processes rate and accelerometer data to form the three-stage missions. The 2.9-m fairing has been proportional and discrete control output commands flight proven over many years, whereas the 3-m needed to drive the control actuators and cold gas jet composite fairing is the new fleet replacement for control thrusters and sequences the remainder of the the standard 3-m aluminum fairing. A new 10L vehicle commands using on-board timing. composite PLF is currently under development. Position and velocity data are explicitly com- It is a 3-m (10-ft) diameter bisector composite puted to derive guidance steering commands. Early fairing whose cylindrical length is .91 m (3 ft) in flight, a load relief mode turns the vehicle into the longer than the baseline 3-m (10-ft) version and wind to reduce angle of attack, structural loads, and with a blunter nose. control effort. After dynamic pressure decay, the The fairings incorporate interior acoustic ab- guidance system corrects trajectory dispersions caused by load relief and directs the vehicle to the sorption blankets as well as flight-proven contami- nominal end-of-stage orbit. Space vehicle separa- nation-free separation joints. McDonnell tion in the desired transfer orbit is accomplished by Douglas provides mission-specific modifications applying time adjustments to the nominal sequence. to the fairings as required by the customer. These include access doors, additional acoustic blankets, 1.3 VEHICLE AXES/ATTITUDE DEFINITIONS and RF windows. Fairings are discussed in greater The axes of the vehicle are defined in Figure 1-6. detail in Section 3. The vehicle centerline is the longitudinal axis of the

1.2.6 Guidance, Control, and Navigation vehicle. Axis II is on the downrange side of the System vehicle and axis IV on the uprange side. The vehicle In the fall of 1995, the Delta II launch vehicles pitches about axes I and III. Positive pitch rotates incorporated the newly developed avionics upgrades the nose of the vehicle up, toward axis IV. The to the Delta Inertial Guidance System (DIGS). The vehicle yaws about axes II and IV. Positive yaw major element of the avionics upgrade is the Redun- rotates the nose of the vehicle to the right, toward dant Inertial Flight Control Assembly (RIFCA) axis I. The vehicle rolls about the centerline. Posi- with its integrated software design, which is a tive roll is clockwise rotation, looking forward (i.e., modernized, single-fault-tolerant guidance system. from axis I toward II). The third-stage spin table RIFCA utilizes six Allied Signal RL20 ring laser also spins in the same direction (i.e., the positive roll gyros and six Sundstrand model QA3000 acceler- direction). ometers to provide redundant three-axis rate and acceleration data. In addition to RIFCA, both the 1.4 LAUNCH VEHICLE INSIGNIA first- and second-stage avionics include a power Delta II users may request a mission-specific and control (P&C) box to support power distribu- insignia to be placed on their launch vehicles. The tion, an ordnance box to issue ordnance commands, user is invited to submit the proposed design to the 1-7 03297REU7

Nose Cone mm in.

2.9-m-dia (9.5-ft) Fairing Air Conditioning Door 8488 2.9-m-dia (9.5-ft) Skin and Stringer Cylinder 334.2 Spacecraft Access Door (as Required)

Contamination-Free Separation Joint

2.4-m-dia (8-ft) Base, Isogrid

Nose Cone

3-m-dia (10-ft) Composite Fairing

Air Conditioning Door 8875 349.4

3-m-dia (10-ft) Cylinder Spacecraft Access Door (as Required)

Second-Stage Access Door (2 Places) Contamination-Free Separation Joint 2.4-m-dia (8-ft) Base

Nose Cone

3-m-dia (10-ft) Fairing10L (Under Development)

Air Conditioning Door 9252.4 364.3 3-m-dia (10-ft) Cylinder

Spacecraft Access Door (as Required)

Contamination-Free Separation Joint

2.4-m-dia (8-ft) Base Figure 1-5. Delta Payload Fairings 1-8 H54536.1 CL

Note: Arrow shows direction of positive vehicle roll

Roll IV CL +XLV Yaw IV III

I +

III

+YLV I Pitch

+ II

+ZLV

Figure 1-6. Vehicle Axes

Delta Program Office no later than 9 months prior Following approval, the Delta Program Office will to launch for review and approval. The maximum have the flight insignia prepared and placed on the size of the insignia is 2.4 by 2.4 m (8 by 8 ft). uprange side of the launch vehicle.

1-9 Section 2 H54963.1 GENERAL PERFORMANCE CAPABILITY

The Delta II can accommodate a wide range of spacecraft requirements. The following sections detail specific performance capabilities of several Delta launch vehicle configurations from the east- Hohmann Transfer ern and western ranges. In addition to the capabili- Restart ties shown herein, our mission designers can pro- SECO 2 vide innovative performance trades to meet the SECO 1 particular requirements of our payload customers. MECO Separation 2.1 LAUNCH SITES Depending on the specific mission requirement Launch and range safety restrictions, the Delta II 7300, 7400 or 7900 series vehicle can make use of either an east or west coast launch site. ■ Eastern Launch Site. The eastern launch site for Delta II is Space Launch Complex 17 (SLC-17) at the Cape Canaveral Air Station (CCAS). This Figure 2-1. Typical Two-Stage Mission Profile site can accommodate flight azimuths in the range of 65 to 110 degrees, with 95 degrees being the most H54964 commonly flown. ■ Western Launch Site. The western launch site for Delta II is Space Launch Complex 2-West (SLC-2W) at the Vandenberg Air Force Base (VAFB). Flight azimuths in the range of 190 to 200 degrees are currently approved by the 30th Space SECO 1 Restart Wing, with 196 degrees being the most commonly SECO 2 flown. Third- MECO Stage Launch 2.2 MISSION PROFILES Burn

Profiles for both two- and three-stage missions Spacecraft are shown in Figures 2-1 and 2-2. Separation ■ 7300 Series Vehicle. In launches from both eastern and western sites, the first stage RS-27A engine and the three strap-on solid-rocket motors are ignited on the ground at liftoff. The solids are then jettisoned following burnout. The main engine continues to burn until main engine cutoff (MECO) at propellant depletion. Figure 2-2. Typical Three-Stage Mission Profile 2-1 ■ 7400 Series Vehicle. MDA will begin flying the In the typical two-stage mission (Figure 2-1), the Delta II with four strap-on solid rocket motors second stage burns for approximately 340 to 420 (Delta 7420) from the eastern launch site in 1997. seconds, at which time second-stage engine cutoff This configuration will be available to spacecraft (SECO 1) occurs. The vehicle then follows a contractors who require more performance than is Hohmann transfer trajectory to the desired low achieved by the 7300 series vehicle: the perfor- Earth orbit (LEO) altitude. Near apogee of the mance capability of the 7400 series vehicle though transfer orbit, the second stage is reignited and not presented, is approximately 11% greater. completes its burn to circularize the orbit. Space- The 7400 series Delta II is available in both two- craft separation takes place approximately 250 sec- and three-stage configurations for launches from onds after second-stage engine cutoff command the eastern and western launch sites. The first-stage (SECO 2). RS-27A engine and the four strap-on solid rocket The three-stage mission, typically a mission to a motors are ignited on the ground at liftoff. The Geosynchronous Transfer Orbit (GTO) solids are jettisoned following burnout, in pairs at a (Figure 2-2), uses the first burn of the second stage one-second interval. The remaining vehicle se- to place the spacecraft into a 185-km (100-nmi) quence of events is approximately the same as with parking orbit inclined at 28.7 degrees. The vehicle the 7300 series vehicle. then coasts to a position near the equator where the ■ 7900 Series Vehicle. In launches from both second stage is restarted and burned until second eastern and western sites, the first-stage RS-27A cutoff. The third stage is spun-up, separated, and burned to establish the GTO. Depending on mis- main engine and six of the nine strap-on solid- sion requirements and spacecraft mass, some incli- rocket motors are ignited on the ground at liftoff. nation can be removed via the burn sequence out of Following burnout of the six solids, the remaining the Earth parking orbit. three GEMs having extended nozzles are ignited. After payload separation, the Delta second stage The six spent cases are then jettisoned in sets of is restarted to deplete any remaining propellants three after vehicle and range safety constraints have (depletion burn) and/or to move the stage to a safe been met. Jettisoning of the second set occurs one distance from the spacecraft (evasive burn). second after the first. The remaining three solids are If required, the multiple restart capability of the jettisoned about three seconds after they burn out. Delta second stage allows the use of an ascending- The main engine then continues to burn until MECO. node flight profile to meet mission requirements. The remainder of the two- and three-stage mis- Typical sequences for LEO missions for the 7320 sion profiles for the 7300 series and 7900 series vehicle from the eastern and western ranges are vehicles are almost identical. Following a short shown in Figures 2-3 and 2-4, while sequences for coast period, separation of the first and second a GTO mission using the 7925 vehicle and a polar stages occurs and, approximately five seconds later, mission using the 7920 vehicle are shown in the second stage ignites. The next major event is the Figures 2-5 and 2-6. Typical event times for both payload fairing (PLF) separation, which occurs two-and three-stage versions of the 7300 series and early in the second-stage flight after an acceptable 7900 series configurations from the eastern and free-molecular heating rate has been reached. western ranges are presented in Tables 2-1 and 2-2. 2-2 02780REU7 SECO I Spacecraft (664 sec) SECO II (3589 sec)

Fairing Drop (300 sec) Restart Stage II Stage II Ignition (3564 sec) Spacecraft (274 sec) Separation (3839 sec) MECO (261 sec)

SRM Drop (3) (66 sec)

Event VI (fps) Acceleration (g)

Liftoff 1343 1.32 Three Solid Motors 3 SRM Burnout 3060 0.94 Burnout (63 sec) MECO 17,162 6.31 SECO I 26,320 0.99 SECO ll 24,084 1.09 Liftoff Main Engine and Three Solid Motors Ignited

Equator

Eastern Range launch site, flight azimuth 95 deg; maximum capability to 28.7-deg inclined orbit, 550-nmi circular

Figure 2-3. Typical Delta II 7320 Mission Profile—Circular Orbit Mission (Eastern Range) (Download Figure)

02781REU7.1 SECO I (666 sec) SECO II (3591 sec) Fairing Drop (290 sec) Restart Stage II Stage II Ignition (3566 sec) Spacecraft (274 sec) Separation (3841 sec) MECO (261 sec)

SRM Drop (3) (110 sec)

Event VI (fps) Acceleration (g)

Liftoff 1255 1.32 Three Solid Motors 3 SRM Burnout 2191 1.00 Burnout (64 sec) MECO 15,847 6.62 SECO I 26,320 1.19 SECO ll 24,084 1.29

Liftoff Main Engine and Three Solid Motors South Pole Ignited

Western Range launch site, flight azimuth 196 deg; maximum capability to polar orbit, 550-nmi circular

Figure 2-4. Typical Delta II 7320 Mission Profile—Polar Orbit Mission (Western Range) (Download Figure) 2-3 02782REU7.1 SECO I SECO II (617 sec) (1306 sec) Stage III Burnout (1483 sec) Fairing Drop Restart Stage II (300 sec) (1237 sec) Stage III Ignition (1396 sec) Stage II Ignition Spacecraft (274 sec) Separation (1596 sec) MECO (261 sec)

SRM Drop (3) Event VI (fps) Acceleration (g) (132 sec) Liftoff 1343 1.37 6 SRM Burnout 3339 0.55 MECO 19,944 5.91 SECO I 25,560 0.67 SRM Drop (6) SECO II 27,192 0.76 (66/67 sec) Stage III Burnout 33,589 3.24

Three Solid Motor Ignition (65.5 sec) Six Solid Motors Burnout (63 sec) Liftoff Main Engine and Six Solid Motors Ignited Equator

Eastern Range launch site, flight azimuth 95 deg; maximum capability to 28.7-deg inclined GTO, 100-nmi perigee

Figure 2-5. Typical Delta II 7925 Mission Profile—GTO Mission (Eastern Range) (Download Figure) 02783REU7.1 SECO I (669 sec) SECO II (3594 sec) Fairing Drop (280 sec) Restart Stage II (3569 sec) Stage II Ignition Spacecraft (274 sec) Separation (3844 sec) MECO (261 sec)

SRM Drop (3) (132 sec)

Event VI (fps) Acceleration (g) SRM Drop (6) (86/87 sec) Liftoff 1255 1.36 6 SRM Burnout 2330 0.59 MECO 18,627 6.22 Three Solid Motor Ignition (66 sec) SECO I 26,320 0.82 SECO II 24,084 0.92 Six Solid Motors Burnout (64 sec) Liftoff Main Engine and Six Solid Motors Ignited

South Pole

Western Range launch site, flight azimuth 196 deg; maximum capability to polar orbit, 550-nmi circular

Figure 2-6. Typical Delta II 7920 Mission Profile—Polar Mission (Western Range) (Download Figure) 2-4 Table 2-1. Delta II Typical Eastern Range Event Times Event 7320 7325 7920 7925 First Stage Main engine ignition T + 0 sec T + 0 T + 0 T + 0 Solid motor ignition (6 solids) T + 0 T + 0 Solid motor burnout (6 solids) T + 63 T + 63 Solid motor ignition (3 solids) T + 0 T + 0 T + 66 T + 66 Solid motor separation (3/3 solids) T + 66/67 T + 66/67 Solid motor burnout (3 solids) T + 63 T + 63 T + 129 T + 129 Solid motor separation (3 solids) T + 66 T + 66 T + 132 T + 132 MECO (M) T + 261 T + 261 T + 261 T + 261 Second Stage Activate stage I/II separation bolts M + 8 M + 8 M + 8 M + 8 Stage II ignition M + 13.5 M + 13.5 M + 13.5 M + 13.5 Fairing separation M + 39 M + 39 M + 39 M + 39 SECO (S1) M + 390 M + 415 M + 408 M + 356 Stage II engine restart S1 + 2900 S1 + 610 S1 + 2900 S1 + 620 SECO (S2) S1 + 2925 S1 + 631 S1 + 2925 S1 + 689 Third Stage Activate spin rockets, start Stage III S2 + 50 S2 + 50 sequencer Separate Stage II S2 + 53 S2 + 53 Stage III ignition S2 + 90 S2 + 90 Stage III burnout S2 + 177 S2 + 177 Spacecraft Spacecraft separation in a 550-nmi circular S2 + 250 S2 + 250 orbit All times shown in seconds. M029,T003, 5/13/96 , 9:44 AM

Table 2-2. Delta II Typical Western Range Event Times Event 7320 7920 7925 First Stage Main engine ignition T + 0 sec T + 0 T + 0 Solid motor ignition (6 solids) T + 0 T + 0 Solid motor burnout (6 solids) T + 64 T + 64 Solid motor ignition (3 solids) T + 0 T + 66 T + 66 Solid motor separation (3/3 solids) T + 86/87 T + 86/87 Solid motor burnout (3 solids) T + 64 T + 130 T + 130 Solid motor separation (3 solids) T + 110 T + 132 T + 132 MECO (M) T + 261 T + 261 T + 261 Second Stage Activate stage I/II separation bolts M + 8 M + 8 M + 8 Stage II ignition M + 13.5 M + 13.5 M + 13.5 Fairing separtion M + 29 M + 19 M + 19 SECO (S1) M + 392 M + 408 M + 356 Stage II engine restart S1 + 2900 S1 + 2900 S1 + 620 SECO (S2) S1 + 2925 S1 + 2925 S1 + 689 Third Stage Activate spin rockets, start stage III S2 + 50 sequencer Separate stage II S2 + 53 Stage III ignition S2 + 90 Stage III burnout S2 + 177 Spacecraft Spacecraft separation in a 550-nmi circular S2 + 250 S2 + 250 orbit All times shown in seconds. M029, T004, 5/13/96, 9:44 AM 2.3 PERFORMANCE CAPABILITY (CCAS) and western (VAFB) launch sites. This section presents a summary of the perfor- The performance estimates shown in mance capabilities of both the two- and three-stage Figures 2-7 through 2-28 that follow are computed vehicles, 7300 and 7900 series, from the eastern based on the following assumptions: 2-5 A. Nominal propulsion system and weight mod- missions. It should be noted that two-stage PAFs els including avionics upgrades were used on all other than those assumed can be used on the 7320 stages. and 7920 configurations. This will affect the space- B. The first stage is burned to propellant craft weight capability shown in the plots that depletion. follow. C. Extended nozzle airlit GEMs are incorporated I. Capabilities are shown for the standard 2.9-m (only airlit GEMs have extended nozzles). (9.5-ft), and 3-m (10-ft) PLFs. D. Second-stage propellant reserve is sufficient A summary of maximum performance for com- to provide a 99.7% probability of a command shut- mon missions is presented in Table 2-3. down (PCS) by the guidance system. Performance data are presented in the following E. PLF separation occurs at a time when the free pages for both two- and three-stage vehicles of the molecular heating rate range is equal to or less than 7300 and 7900 series launched from the eastern and 1135 W/m2 ( 0.1 Btu/ft2 sec). western ranges. Spacecraft weight capability is F. Perigee velocity is the vehicle burnout veloc- presented as a function of the parameters listed ity at 185 km (100 nmi) altitude and zero degree below. flight path angle. G. Initial flight azimuth is 95 degrees from the 7300 SERIES VEHICLE. eastern launch site, and 196 degrees from the west- ■ Eastern Launch Site. ern launch site. - Two-stage perigee velocity (Figure 2-7). H. A 6019 payload attach fitting (PAF) is as- - Two-stage apogee altitude (Figure 2-8). sumed for 7300 series two-stage missions, and a - Two-stage circular orbit altitude (Figure 2-9). 6915 PAF is assumed for 7900 series two-stage - Three-stage perigee velocity (Figure 2-10). missions, while a 3712 A, B or C PAF is assumed - Three-stage planetary mission launch energy for both 7300 series and 7900 series three-stage (Figure 2-11).

Table 2-3. Mission Capabilities Spacecraft Weight (kg/lb) 7325 7325-10 7925 7925-10 Three-stage missions ■ Geosynchronous Transfer Orbit (GTO) **1869/ 1798/ ● CCAS, i = 28.7 deg 4120 3965 ● 185 x 35,786 km/100 x 19,323 nmi ■ Molniya orbit ● VAFB, i = 63.4 deg N/A N/A 1220/ 1170/ ● 370 x 40,094 km/200 x 21,649 nmi 2690 2580 7320 7320-10 7920 7920-10 Two-stage missions ■ Low Earth orbit (LEO) 2867/ 2735/ 5139/ 4971/ ● CCAS, i = 28.7 deg 6320 6030 11,330 10,960 ● 185 km/100 nmi circular ■ LEO ● VAFB, i = 90 deg 2096/ 1982 / 3896/ 3774/ ● 185 km/100 nmi circular 4620 4370 8590 8320 ■ Sun-synchronous orbit ● VAFB, i = 98.7 deg 1687/ 1578 3220/ 3112/ ● 833 km/450 nmi circular 3720 3480 7100 6860 *Requires an impacting second-stage flight mode. Direct inquiries to the Delta Program Office M029, T005, 12/11/95 , 10 :16 AM 2-6 ■ Western Launch Site. 2.4 MISSION ACCURACY DATA - Two-stage perigee velocity (Figure 2-12). All Delta II configurations employ the RIFCA - Two-stage apogee altitude (Figure 2-13). mounted in the second-stage guidance compartment. - Two-stage circular orbit altitude This system provides precise pointing and orbit (Figure 2-14). accuracy for both two- and three-stage missions. - Two-stage sun-synchronous orbit The typical two-stage mission to low-altitude (Figure 2-15). circular orbit uses a Hohmann transfer between

7900 Series Vehicle. second stage burns in which the second stage is restarted to circularize the orbit after a coast period. ■ Eastern Launch Site. In these cases the three-sigma dispersion accuracy – Two-stage perigee velocity (Figure 2-16). achieved by the guidance system is typically less – Two-stage apogee altitude (Figure 2-17). than the following: – Two-stage circular orbit altitude (Figure 2-18). ■ ± ± – Three-stage perigee velocity (Figure 2-19). Circular orbit altitude: 9.3 km ( 5 nmi). ± – Three-stage apogee altitude (Figure 2-20). ■ Orbit inclination: 0.05 degrees. – Three-stage GTO inclination (Figure 2-21). In a three-stage mission, the parking-orbit pa- – Three-stage planetary mission rameters achieved are quite accurate. The final (Figure 2-22). orbit (e.g., GTO) is primarily affected by the third ■ Western Launch Site. stage pointing and the impulse errors from the third – Two-stage perigee velocity (Figure 2-23). stage solid motor burn. The pointing error for a – Two-stage apogee altitude (Figure 2-24). given mission depends on the third-stage/space- – Two-stage circular orbit altitude craft mass properties and the spin rate. The typical (Figure 2-25). pointing error at third stage ignition is approxi- – Two-stage sun-synchronous orbit (Figure 2-26). mately 1.5 degrees based on past Delta experience. – Three-stage perigee velocity (Figure 2-27). Apogee altitude variations for the GTO mission that – Three-stage apogee altitude (Figure 2-28). result from this error are shown in Figure 2-29. The The performance capability for any given mis- transfer orbit inclination error is typically from ±0.2 sion depends upon quantitative analysis of all to ±0.6 degrees over the range shown, while the known mission requirements and range safety re- perigee altitude variation is typically about ±5.6 km strictions. The allowable spacecraft weight should (±3 nmi). All errors are three-sigma values. be coordinated with the Delta Program Office as early as possible in the basic mission planning. These data are presented as general indicators Preliminary error analysis, performance optimiza- only. Individual mission requirements and specifi- tion, and trade-off studies will be performed, as cations will be used to perform detailed analyses for required, to arrive at an early commitment of allow- specific missions. The user is invited to contact the able spacecraft weight for each specific mission. Delta Program Office for further information.

2-7 02784REU7.1

37,000

36,000 Legend 9.5-ft Fairing 35,000 10-ft Fairing

34,000

33,000

32,000

31,000

30,000

Perigee Velocity (ft/sec) 29,000

28,000

95-deg Flight Azimuth 27,000 28.7-deg Inclination 100-nmi Perigee Altitude 26,000 155-lb Payload Attach Fitting

25,000 0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000 5500 6000 6500 7000 Spacecraft Weight (lb)

11.5

Legend 11.0 2.9-m Fairing 3.0-m Fairing

10.5

10.0

9.5

9.0 Perigee Velocity (km/sec)

8.5

95-deg Flight Azimuth 28.7-deg Inclination 8.0 185-km Perigee Altitude 70.3-kg Payload Attach Fitting

7.5 0 200 400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 2600 2800 3000 3200 Spacecraft Mass (kg)

Figure 2-7. Delta II 7320 Vehicle, Two-Stage Perigee Velocity Capability, Eastern Range (Download Figure) 2-8 02786REU7.1

24,000

22,000 Legend 9.5-ft Fairing 20,000 10-ft Fairing

18,000

16,000

14,000

12,000

10,000 Apogee Altitude (nmi) 8000

6000

4000 95-deg Flight Azimuth 28.7-deg Inclination 100-nmi Perigee Altitude 2000 155-lb Payload Attach Fitting

0 0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000 5500 6000 6500 7000 Spacecraft Weight (lb)

45,000

Legend 40,000 2.9-m Fairing 3.0-m Fairing

35,000

30,000

25,000

20,000 Apogee Altitude (km) 15,000

10,000 95-deg Flight Azimuth 28.7-deg Inclination 5000 185-km Perigee Altitude 70.3-kg Payload Attach Fitting 0 0 200 400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 2600 2800 3000 3200 Spacecraft Mass (kg)

Figure 2-8. Delta II 7320 Vehicle, Two-Stage Apogee Altitude Capability, Eastern Range (Download Figure) 2-9 02787REU7.1

6000

5500 Legend 9.5-ft Fairing 5000 10-ft Fairing

4500

4000

3500

3000

2500

Circular Orbit Altitude (nmi) 2000

1500

1000 95-deg Flight Azimuth 28.7-deg Inclination 500 155-lb Payload Attach Fitting

0 0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000 5500 6000 6500 7000 Spacecraft Weight (lb)

11,000

10,000 Legend 2.9-m Fairing 9000 3.0-m Fairing

8000

7000

6000

5000

4000 Circular Orbit Altitude (km)

3000

2000 95-deg Flight Azimuth 28.7-deg Inclination 1000 70.3-kg Payload Attach Fitting

0 0 200 400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 2600 2800 3000 3200 Spacecraft Mass (kg)

Figure 2-9. Delta II 7320 Vehicle, Two-Stage Circular Orbit Capability, Eastern Range (Download Figure) 2-10 02788REU7.1

47,000

46,000 Legend 9.5-ft Fairing 45,000 10-ft Fairing 44,000 Star-48B Offload Max Star-48B Offload 43,000

42,000 Note: Spacecraft weights less than 1500 lb 41,000 may require nutation control system modifications which may result in a decrease in spacecraft weight 40,000

39,000

Perigee Velocity (ft/sec) 38,000

37,000

36,000 95-deg Flight Azimuth 35,000 28.7-deg Inclination 100-nmi Perigee Altitude 34,000 100-lb Payload Attach Fitting

33,000 0 200 400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 Spacecraft Weight (lb)

14.5

Legend Legend 14.0 2.9-m Fairing 3.0-m Fairing Star-48B Offload 13.5 Max Star-48B Offload

13.0 Note: Spacecraft mass less than 680 kg may require nutation control system 12.5 modifications which may result in a decrease in spacecraft mass 12.0

Perigee Velocity (km/sec) 11.5

11.0 95-deg Flight Azimuth 28.7-deg Inclination 10.5 185-km Perigee Altitude 45.4-kg Payload Attach Fitting

10.0 0 100 200 300 400 500 600 700 800 900 1000 1100 Spacecraft Mass (kg)

Figure 2-10. Delta II 7325 Vehicle, Three-Stage Perigee Velocity Capability, Eastern Range (Download Figure) 2-11 02789REU7.1

1800

1600 Legend 9.5-ft Fairing 10-ft Fairing 1400 Star-48B Offload

1200 Note: Spacecraft weights less than 1500 lb may require nutation control system 1000 modifications which may result in a decrease in spacecraft weight

800 Spacecraft Weight (lb) 600

400 95-deg Flight Azimuth 28.7-deg Inclination 200 100-nmi Perigee Altitude 100-lb Payload Attach Fitting

0 0 5 10 15 20 25 30 35 40 45 50 55 60 Launch Energy (km2/sec2)

800

700 Legend 2.9-m Fairing 3.0-m Fairing 600 Star-48B Offload

500 Note: Spacecraft mass less than 680 kg may require nutation control system modifications which may result in a 400 decrease in spacecraft mass

300 Spacecraft Mass (kg)

200

95-deg Flight Azimuth 28.7-deg Inclination 100 185-km Perigee Altitude 45.4-kg Payload Attach Fitting

0 0 5 10 15 20 25 30 35 40 45 5055 60 Launch Energy (km2/sec2)

Figure 2-11. Delta II 7325 Vehicle, Three-Stage Planetary Mission Capability, Eastern Range (Download Figure) 2-12 02790REU7.1

36,000

35,000 Legend 9.5-ft Fairing 34,000 10-ft Fairing

33,000

32,000

31,000

30,000

29,000 Perigee Velocity (ft/sec)

28,000

27,000 196-deg Flight Azimuth 90-deg Inclination 100-nmi Perigee Altitude 26,000 155-lb Payload Attach Fitting

25,000 0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000 5500 Spacecraft Weight (lb)

11.0

Legend 10.5 2.9-m Fairing 3.0-m Fairing

10.0

9.5

9.0 Perigee Velocity (km/sec)

8.5

196-deg Flight Azimuth 90-deg Inclination 8.0 185-km Perigee Altitude 70.3-kg Payload Attach Fitting

7.5 0 200 400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 Spacecraft Mass (kg)

Figure 2-12. Delta II 7320 Vehicle, Two-Stage Perigee Velocity Capability, Western Range (Download Figure) 2-13 02791REU7.1

24,000

22,000 Legend 9.5-ft Fairing 20,000 10-ft Fairing

18,000

16,000

14,000

12,000

10,000 Apogee Altitude (nmi) 8000

6000

4000 196-deg Flight Azimuth 90-deg Inclination 2000 100-nmi Perigee Altitude 155-lb Payload Attach Fitting 0 0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000 5500 Spacecraft Weight (lb)

45,000

Legend 40,000 2.9-m Fairing 3.0-m Fairing

35,000

30,000

25,000

20,000

Apogee Altitude (km) 15,000

10,000 196-deg Flight Azimuth 90-deg Inclination 5000 185-km Perigee Altitude 70.3-kg Payload Attach Fitting 0 0 200 400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 Spacecraft Mass (kg)

Figure 2-13. Delta II 7320 Vehicle, Two-Stage Apogee Altitude Capability, Western Range (Download Figure) 2-14 02792REU7.1

6000

5500 Legend 9.5-ft Fairing 5000 10-ft Fairing

4500

4000

3500

3000

2500

2000 Circular Orbit Altitude (nmi)

1500

1000 196-deg Flight Azimuth 90-deg Inclination 500 155-lb Payload Attach Fitting

0 0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000 5500 Spacecraft Weight (lb)

11,000

10,000 Legend 2.9-m Fairing 9000 3.0-m Fairing

8000

7000

6000

5000

4000 Circular Orbit Altitude (km)

3000

2000 196-deg Flight Azimuth 90-deg Inclination 1000 70.3-kg Payload Attach Fitting

0 0 200 400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 Spacecraft Mass (kg)

Figure 2-14. Delta II 7320 Vehicle, Two-Stage Circular Orbit Capability, Western Range (Download Figure) 2-15 02793REU7.1

2400

2200 Legend 9.5-ft Fairing 2000 10-ft Fairing

1800

1600

1400

1200

1000

800 Circular Orbit Altitude (nmi)

600

400 196-deg Minimum Flight Azimuth Variable Inclination 200 155-lb Payload Attach Fitting

0 0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000 5500 Spacecraft Weight (lb)

4400

4000 Legend 2.9-m Fairing 3600 3.0-m Fairing

3200

2800

2400

2000

1600 Circular Orbit Altitude (km)

1200

800 196-deg Minimum Flight Azimuth Variable Inclination 400 70.3-kg Payload Attach Fitting

0 0 200 400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 Spacecraft Mass (kg)

Figure 2-15. Delta II 7320 Vehicle, Two-Stage Sun-Synchronous Orbit Capability, Western Range (Download Figure) 2-16 02794REU7.1

38,000

37,000 Legend 9.5-ft Fairing 36,000 10-ft Fairing

35,000

34,000

33,000

32,000

31,000

30,000 Perigee Velocity (ft/sec) 29,000

28,000 95-deg Flight Azimuth 27,000 28.7-deg Inclination 100-nmi Perigee Altitude 26,000 225-lb Payload Attach Fitting

25,000 0 1000 2000 3000 4000 5000 6000 7000 8000 9000 10,00011,000 12,000 Spacecraft Weight (lb)

12.0

Legend 11.5 2.9-m Fairing 3.0-m Fairing 11.0

10.5

10.0

9.5

Perigee Velocity (km/sec) 9.0

8.5 95-deg Flight Azimuth 28.7-deg Inclination 8.0 185-km Perigee Altitude 102.1-kg Payload Attach Fitting

7.5 0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000 5500 Spacecraft Mass (kg)

Figure 2-16. Delta II 7920 Vehicle, Two-Stage Perigee Velocity Capability, Eastern Range (Download Figure) 2-17 02795REU7.1

24,000

22,000 Legend 9.5-ft Fairing 20,000 10-ft Fairing

18,000

16,000

14,000

12,000

10,000 Apogee Altitude (nmi) 8000

6000

4000 95-deg Flight Azimuth 28.7-deg Inclination 100-nmi Perigee Altitude 2000 225-lb Payload Attach Fitting

0 0 1000 2000 3000 4000 5000 6000 7000 8000 9000 10,00011,000 12,000 Spacecraft Weight (lb)

45,000

Legend 40,000 2.9-m Fairing 3.0-m Fairing

35,000

30,000

25,000

20,000 Apogee Altitude (km) 15,000

10,000 95-deg Flight Azimuth 28.7-deg Inclination 5000 185-km Perigee Altitude 102.1-kg Payload Attach Fitting

0 0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000 5500 Spacecraft Mass (kg)

Figure 2-17. Delta II 7920 Vehicle, Two-Stage Apogee Altitiude Capability, Eastern Range (Download Figure) 2-18 02796REU7.1

6000

5500 Legend 9.5-ft Fairing 5000 10-ft Fairing

4500

4000

3500

3000

2500

Circular Orbit Altitude (nmi) 2000

1500

1000 95-deg Flight Azimuth 500 28.7-deg Inclination 225-lb Payload Attach Fitting 0 0 1000 2000 3000 4000 5000 6000 7000 8000 9000 10,00011,000 12,000 Spacecraft Weight (lb)

11,000

10,000 Legend 2.9-m Fairing 9000 3.0-m Fairing

8000

7000

6000

5000

4000 Circular Orbit Altitude (km)

3000

2000 95-deg Flight Azimuth 1000 28.7-deg Inclination 102.1-kg Payload Attach Fitting

0 0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000 5500 Spacecraft Mass (kg)

Figure 2-18. Delta II 7920 Vehicle, Two-Stage Circular Orbit Capability, Eastern Range (Download Figure) 2-19 02797REU7.1

48,000

Legend 46,000 9.5-ft Fairing 10-ft Fairing

44,000

42,000 Note: Spacecraft weights less than 1500 lb may require nutation control system 40,000 modifications which may result in a decrease in spacecraft weight

38,000

Perigee Velocity (ft/sec) 36,000

34,000 95-deg Flight Azimuth 28.7-deg Inclination 32,000 100-nmi Perigee Altitude 100-lb Payload Attach Fitting

30,000 0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000 5500 Spacecraft Weight (lb)

14.5

14.0 Legend 2.9-m Fairing 13.5 3.0-m Fairing

13.0

12.5 Note: Spacecraft mass less than 680 kg 12.0 may require nutation control system modifications which may result in a 11.5 decrease in spacecraft mass

11.0

Perigee Velocity (km/sec) 10.5

10.0

95-deg Flight Azimuth 9.5 28.7-deg Inclination 185-km Perigee Altitude 9.0 45.4-kg Payload Attach Fitting

8.5 0 200 400 600 800 1000 1200 1400 1600 1800 20002200 2400 Spacecraft Mass (kg)

Figure 2-19. Delta II 7925 Vehicle, Three-Stage Perigee Velocity Capability, Eastern Range (Download Figure) 2-20 02798REU7.1

120,000

Legend 9.5-ft Fairing 100,000 10-ft Fairing

80,000

60,000 Apogee Altitude (nmi) 40,000

95-deg Flight Azimuth 20,000 28.7-deg Inclination 100-nmi Perigee Altitude 100-lb Payload Attach Fitting

0 0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000 5500 Spacecraft Weight (lb)

200,000

180,000 Legend 2.9-m Fairing 3.0-m Fairing 160,000

140,000

120,000

100,000

80,000 Apogee Altitude (km)

60,000

40,000 95-deg Flight Azimuth 28.7-deg Inclination 185-km Perigee Altitude 20,000 45.4-kg Payload Attach Fitting

0 0 200 400 600 800 1000 1200 1400 1600 1800 20002200 2400 Spacecraft Mass (kg)

Figure 2-20. Delta II 7925 Vehicle, Three-Stage Apogee Altitude Capability, Eastern Range (Download Figure) 2-21 02799REU7

4300

4200 Legend 9.5-ft Fairing 4100 10-ft Fairing

4000

3900

3800

3700

Spacecraft Weight (lb) 3600

3500

3400 95-deg Flight Azimuth 100-nmi Perigee Altitude 3300 100-lb Payload Attach Fitting

3200 20 21 22 23 24 25 26 27 28 29 30 GTO Inclination (deg)

1950

1900 Legend 2.9-m Fairing 1850 3.0-m Fairing

1800

1750

1700

1650

Spacecraft Mass (kg) 1600

1550

1500 95-deg Flight Azimuth 185-km Perigee Altitude 1450 45.4-kg Payload Attach Fitting

1400 20 21 22 23 24 25 26 27 28 29 30 GTO Inclination (deg)

Figure 2-21. Delta II 7925 Vehicle, Three-Stage GTO Inclination Capability, Eastern Range (Download Figure) 2-22 02800REU7.1

3000

Legend 9.5-ft Fairing 2500 10-ft Fairing

Note: Spacecraft weights less than 1500 lb 2000 may require nutation control system modifications which may result in a decrease in spacecraft weight

1500 Spacecraft Weight (lb) 1000

95-deg Flight Azimuth 500 28.7-deg Inclination 100-nmi Perigee Altitude 100-lb Payload Attach Fitting

0 0 5 10 15 20 25 30 35 40 45 5055 60 Launch Energy (km2/sec2)

1400

Legend 1200 2.9-m Fairing 3.0-m Fairing

1000 Note: Spacecraft mass less than 680 kg may require nutation control system modifications which may result in a decrease in spacecraft mass 800

600 Spacecraft Mass (kg)

400

95-deg Flight Azimuth 200 28.7-deg Inclination 185-km Perigee Altitude 45.4-kg Payload Attach Fitting

0 0 5 10 15 20 25 30 35 40 45 5055 60 Launch Energy (km2/sec2)

Figure 2-22. Delta II 7925 Vehicle, Three-Stage Planetary Mission Capability, Eastern Range (Download Figure) 2-23 02801REU7.1

37,000

36,000 Legend 9.5-ft Fairing 35,000 10-ft Fairing

34,000

33,000

32,000

31,000

30,000

Perigee Velocity (ft/sec) 29,000

28,000

196-deg Flight Azimuth 27,000 90-deg Inclination 100-nmi Perigee Altitude 26,000 225-lb Payload Attach Fitting

25,000 0 1000 2000 3000 4000 5000 6000 70008000 9000 Spacecraft Weight (lb)

11.5

Legend 11.0 2.9-m Fairing 3.0-m Fairing

10.5

10.0

9.5

9.0 Perigee Velocity (km/sec)

8.5 196-deg Flight Azimuth 90-deg Inclination 8.0 185-km Perigee Altitude 102.1-kg Payload Attach Fitting

7.5 0 500 1000 1500 2000 2500 3000 3500 4000 4500 Spacecraft Mass (kg)

Figure 2-23. Delta II 7920 Vehicle, Two-Stage Perigee Velocity Capability, Western Range (Download Figure) 2-24 02785REU7.1

24,000

22,000 Legend 9.5-ft Fairing 20,000 10-ft Fairing

18,000

16,000

14,000

12,000

10,000 Apogee Altitude (nmi) 8000

6000

196-deg Flight Azimuth 4000 90-deg Inclination 100-nmi Perigee Altitude 2000 225-lb Payload Attach Fitting

0 0 1000 2000 3000 4000 5000 6000 70008000 9000 Spacecraft Weight (lb)

45,000

Legend 40,000 2.9-m Fairing 3.0-m Fairing

35,000

30,000

25,000

20,000 Apogee Altitude (km) 15,000

10,000 196-deg Flight Azimuth 90-deg Inclination 5000 185-km Perigee Altitude 102.1-kg Payload Attach Fitting

0 0 500 1000 1500 2000 2500 3000 3500 4000 4500 Spacecraft Mass (kg)

Figure 2-24. Delta II 7920 Vehicle, Two-Stage Apogee Altitude Capability, Western Range (Download Figure) 2-25 02830REU7.1

6000

5500 Legend 9.5-ft Fairing 5000 10-ft Fairing

4500

4000

3500

3000

2500

Circular Orbit Altitude (nmi) 2000

1500

1000 196-deg Flight Azimuth 90-deg Inclination 500 225-lb Payload Attach Fitting

0 0 1000 2000 3000 4000 5000 6000 70008000 9000 Spacecraft Weight (lb)

11,000

10,000 Legend 2.9-m Fairing 9000 3.0-m Fairing

8000

7000

6000

5000

4000 Circular Orbit Altitude (km)

3000

2000 196-deg Flight Azimuth 90-deg Inclination 1000 102.1-kg Payload Attach Fitting

0 0 500 1000 1500 2000 2500 3000 3500 4000 4500 Spacecraft Mass (kg)

Figure 2-25. Delta II 7920 Vehicle, Two-Stage Circular Orbit Capability, Western Range (Download Figure) 2-26 02831REU7.1

2400

2200 Legend 9.5-ft Fairing 2000 10-ft Fairing

1800

1600

1400

1200

1000

Circular Orbit Altitude (nmi) 800

600

400 196-deg Minimum Flight Azimuth Variable Inclination 200 225-lb Payload Attach Fitting

0 0 1000 2000 3000 4000 5000 6000 70008000 9000 Spacecraft Weight (lb)

4400

4000 Legend 2.9-m Fairing 3600 3.0-m Fairing

3200

2800

2400

2000

1600 Circular Orbit Altitude (km)

1200

800 196-deg Minimum Flight Azimuth Variable Inclination 400 102.1-kg Payload Attach Fitting

0 0 500 1000 1500 2000 2500 3000 3500 4000 Spacecraft Mass (kg)

Figure 2-26. Delta II 7920 Vehicle, Two-Stage Sun-Synchronous Orbit Capability, Western Range (Download Figure) 2-27 02832REU7.1

46,000

Legend 44,000 9.5-ft Fairing 10-ft Fairing Star-48B Offload 42,000

Note: Spacecraft weights less than 1500 lb 40,000 may require nutation control system modifications which may result in a decrease in spacecraft weight 38,000

36,000 Perigee Velocity (ft/sec)

34,000

196-deg Flight Azimuth 90-deg Inclination 32,000 100-nmi Perigee Altitude 100-lb Payload Attach Fitting

30,000 0 500 1000 1500 2000 2500 3000 35004000 4500 Spacecraft Weight (lb)

14.0

13.5 Legend 2.9-m Fairing 3.0-m Fairing 13.0 Star-48B Offload

12.5

Note: Spacecraft mass less than 680 kg 12.0 may require nutation control system modifications which may result in a decrease in spacecraft mass 11.5

11.0 Perigee Velocity (km/sec)

10.5 196-deg Flight Azimuth 90-deg Inclination 10.0 185-km Perigee Altitude 45.4-kg Payload Attach Fitting

9.5 0 200 400 600 800 10001200 1400 1600 18002000 2200 Spacecraft Mass (kg)

Figure 2-27. Delta II 7925 Vehicle, Three-Stage Perigee Velocity Capability, Western Range (Download Figure) 2-28 02833REU7.1

120,000

Legend 9.5-ft Fairing 100,000 10-ft Fairing Star-48B Offload

80,000

Note: Spacecraft weights less than 1500 lb 60,000 may require nutation control system modifications which may result in a decrease in spacecraft weight Apogee Alititude (nmi) 40,000

196-deg Flight Azimuth 20,000 90-deg Inclination 100-nmi Perigee Altitude 100-lb Payload Attach Fitting

0 0 500 1000 1500 2000 2500 3000 35004000 4500 Spacecraft Weight (lb)

200,000

180,000 Legend 2.9-m Fairing 3.0-m Fairing 160,000 Star-48B Offload

140,000

120,000

100,000 Note: Spacecraft mass less than 680 kg may require nutation control system 80,000 modifications which may result in a

Apogee Altitude (km) decrease in spacecraft mass

60,000

40,000 196-deg Flight Azimuth 90-deg Inclination 185-km Perigee Altitude 20,000 45.4-kg Payload Attach Fitting

0 0 200 400 600 800 10001200 1400 1600 1800 2000 Spacecraft Mass (kg)

Figure 2-28. Delta II 7925 Vehicle, Three-Stage Apogee Altitude Capability, Western Range (Download Figure) 2-29 02834REU7

1200

1100

1000

900

800

700

600

500

400 Pointing Error = 1.50 deg Specific Impulse Error = 0.36%

Deviation from Nominal Apogee Altitude, +3 Sigma (nmi) 300

200 20 21 22 23 24 25 26 2728 29 30 Inclination (deg)

2200

2000

1800

1600

1400

1200

1000

800 Pointing Error = 1.50 deg Specific Impulse Error = 0.36% Deviation from Nominal Apogee Altitude, +3 Sigma (km) 600

400 20 21 22 23 24 2526 27 28 29 30 Inclination (deg)

Figure 2-29. Delta II 7925 Vehicle, Three-Stage GTO Deviations Capability, Eastern Range (Download Figure) 2-30 Section 3 included are the manufacturing tolerances of the SPACECRAFT FAIRINGS launch vehicle as well as the thickness of the The spacecraft is protected by a fairing which acoustic blanket installed on the fairing interior shields it from aerodynamic buffeting and heating (including billowing). The blanket configurations while in the lower atmosphere. The fairing is jetti- available are described in Table 3-1. soned during second-stage powered flight at an Clearance layouts and analyses are performed altitude of at least 125 km (67 nmi). A general and, if necessary, critical clearances are measured discussion of the Delta fairings is presented in after the fairing is installed to ensure positive clear- Section 3.1. Detailed descriptions and envelopes ance during flight. To accomplish this, it is impor- for the 2.9-m (9.5-ft) and 3-m (10-ft) fairings are tant that the spacecraft description (refer to presented in Sections 3.2, 3.3, and 3.4 . Section 8) include an accurate definition of the physical location of all points on the spacecraft that 3.1 GENERAL DESCRIPTION are within 51 mm (2 in.) of the allowable envelope. The envelopes presented in the following sec- The dimensions must include the maximum manu- tions define the maximum allowable static dimen- facturing tolerances. sions of the spacecraft (including manufacturing Electrical umbilical cabling to the spacecraft (if tolerances) for the spacecraft/attach fitting inter- required) is attached to the inside surface of the face. If dimensions are maintained within these fairing shell. Electrical disconnect is accomplished envelopes, there will be no contact of the spacecraft at fairing separation by quick-disconnect connectors. with the fairing during flight, provided that the An air-conditioning inlet umbilical door on the frequency and structural stiffness characteristics of fairing provides a controlled environment to the the spacecraft are in accordance with the limits spacecraft and launch vehicle second stage while specified in Section 4.2.3. These envelopes include on the launch stand. A GN2 purge system can be allowances for relative static/dynamic deflections incorporated to provide continuous dry nitrogen to between the launch vehicle and spacecraft. Also the spacecraft until liftoff.

Table 3-1. Typical Acoustic Blanket Configurations Fairing Location 2.9 m (9.5 ft) dia Blankets extend from the nose cap to approximately Station 491. The blanket thicknesses are as follows: 38.1 mm (1.5 in.) in the nose section, 76.2 mm (3.0 in.) in the 2896-mm (114-in.) diameter section, and 38.1 mm (1.5 in.) in the upper portion of the 2438-mm (96-in.) diameter section. 3 m (10 ft) dia The baseline configuration for acoustic blankets extends from the aft end of the boattail to station 208.37 8.8m (29 ft) in the nose section. These blankets are 76.2 mm (3 in.) thick throughout this region. length 3 m (10 ft) dia The baseline configuration for acoustic blankets extends from the aft end of the boattail to station 194.72 9.3m (30.4 ft) in the nose section. These blankets are 76.2 mm (3 in.) thick throughout this region. length ■ These configurations may be modified to meet mission-specific requirements. ■ Blankets for the 9.5-ft Delta fairing are constructed of silicone-bonded heat-treated glass-fiber batt enclosed between two 0.076-mm (0.003-in.) conductive Teflon-impregnated fiberglass facesheets. The blankets are vented through a 5-micron stainless steel mesh filter, which controls particulate contamination to levels better than a Class 10,000 cleanroom environment. ■ Blankets for the 10-ft Delta composite fairings are constructed of acoustic material. The blankets are vented through the interstage. ■ Outgassing of the acoustic blankets meets the criteria of 1.0% maximum total weight loss and 0.10% maximum volatile condensable material for the 2.9m (9.5 ft) fairing. ■ The acoustic blankets for the 3m (10 ft) composite fairings are being designed to meet the intent of the criteria of 1.0% maximum total weight loss and 0.10% maximum volatile condensible material.

M029, T010, 5 /16/96, 02:08 PM 3-1 Contamination of the spacecraft is minimized by rail retains the detonating-fuse gases to prevent factory cleaning prior to shipment to the field site. contamination of the spacecraft during the fairing Special cleaning of the fairing in the field in a clean- separation event. room environment using “black light” is available Two 18-in by 18-in access doors for second- upon request. stage access are part of the baseline fairing configuration (Figure 3-2). To satisfy spacecraft require- 3.2 THE 2.9-m (9.5-ft) DIAMETER SPACECRAFT FAIRING ments, additional removable doors of various sizes and locations can be provided to permit access to The 2.9-m (9.5-ft) Delta fairing (Figures 3-1 and the spacecraft following fairing installation. It should 3-2) is an aluminum structure fabricated in two be noted that the access doors will have acoustic half-shells. These shells consist of a hemispherical blankets. The quantity and location of access doors nose cap, a biconic section, a cylindrical center must also be coordinated with the Delta Program section of 2896-mm (114-in.) diameter (the maxi- Office. mum diameter of the fairing), a 30-degree conical The fiberglass biconic section can be made transition, and a cylindrical base section having the RF-transparent by removal of its aluminum foil 2438-mm (96-in.) core vehicle diameter. The biconic lining. Location and size of the RF panels must be section is a ring-stiffened monocoque structure, coordinated with the Delta Program Office. one-half of which is fiberglass covered with a Acoustic absorption blankets are provided within removable aluminum foil lining to create an RF the fairing interior. The typical blanket configura- window. The cylindrical base section is an inte- tion is described in Table 3-1. Blanket thermal grally stiffened isogrid structure, and the cylindri- characteristics are discussed in Section 4.2.2. cal center section has a skin-and-stringer-construc- The allowable static spacecraft envelopes for tion. The fairing has an overall length of 8488 mm existing attach fittings within the fairing are shown (334.2 in.). in Figures 3-3 and 3-4 and assume that the space- The half-shells are joined by a contamination-free craft stiffness recommended in Section 4.2.3.2 is linear piston/cylinder thrusting separation system maintained. Usable envelopes below the separation that runs longitudinally the full length of the fairing. plane and local protuberances outside the enve- Two functionally redundant explosive bolt assem- lopes presented require coordination and approval blies provide the structural continuity at the base of the Delta Program Office. ring of the fairing. Four functionally redundant explosive bolt assemblies (two each) provide cir- 3.3 THE 3-m (10-ft) DIAMETER cumferential structural continuity at the 30-degree SPACECRAFT FAIRING transition section between the 2896-mm (114-in.) The 3-m (10-ft) diameter fairing is available for diameter section and the 2438-mm (96-in.) diam- spacecraft requiring a larger envelope. The fairing eter section. (Figure 3-5) is a composite sandwich structure The fairing half-shells are jettisoned by actuation which separates into bisectors. Each bisector is of the base and transition separation nuts and by the constructed in a single co-cured lay-up, eliminating detonating fuse in the thrusting joint cylinder rail the need for module-to-module manufacturing joints cavity. A bellows assembly within each cylinder and intermediate ring stiffeners. The resulting 3-2 H30379.3

DAC 116907

Figure 3-1. Delta 2.9-m (9.5-ft) Payload Fairing 3-3 02620REU7.1 mm 676 All dimensions are in R in. 26.60 All station numbers are in inches Sta No. 219.22

RF Transparent 1/2 Nose Section Two-Strand 1563 Detonating 20 deg Fuse Rivets 61.5

299.98 15 deg 853 33.59

Bellows 333.5 A/C Inlet Door Detail B (See Figure 4-1) 356.9 2032 80.00

Explosive Bolt (6 Places) 413.5 30 deg 429.1

Fairing Split Line 396 15.59

8488 Sta 491 3157 334.17 124.29 RIFCA Line-of-Sight Door

457 X 457 519.9 18 X 18 Access Door Base Cylinder (2 Places) 553.39 A A

IV Fairing Split Line 2438 96.00 dia Base Cylinder

lll l 43° 52' CL Access Door 43° 23' Base Cylinder

2896 dia 114 B CL A/C Door Center Cylinder ll CL Access Door View A-A Base Cylinder Figure 3-2. Profile, 2.9-m (9.5-ft) Diameter Fairing (Download Figure) 3-4

02737REU7.1

Fairing Envelope Usable Payload Envelope (2)

Usable Envelope Below Separation Plane Attach Fitting Motor

Sta 219.22 Notes: 1. All dimensions are in mm 523 in. R 20.60 2. All station numbers are in inches 3. Acoustic blanket thickness is 38.1 mm (1.5 in.) 4678 in the nose and 76.2 mm (3 in.) in the cylindrical 184.25 section 4. MDA requires definition of spacecraft features within 50.8mm (2.0 in.) of payload envelope 20 deg 5. Projections of spacecraft appendages below the spacecraft separation plane may be permitted but must be coordinated with the Delta Program 733 Office 28.84 15 deg

2896 dia 114.00

2540 dia 2004 100.00 78.90

2482 97.7 dia

2540 dia 100.00

1243 dia 48.93 Sta 413.95 Spacecraft 940 dia Separation 37.00 Plane for 102 3712 PAF 724 dia 4.00 51 28.50 1.99 8488

334.17

647 73 R 345 25.49 2.88 13.60

30 15 15 deg deg deg

Sta 413.95 Spacecraft Sta 553.39 Separation Plane 2438 dia 96.00

Figure 3-3. Spacecraft Envelope, 2-9-m (9.5-ft) Diameter Fairing, Three-Stage Configuration (3712 PAF) (Download Figure) 3-5 02880REU7 Fairing Envelope

Usable Payload Envelope

Attach Fitting Sta 219.22 Notes: mm 1. All dimensions are in 523 in. R 2. All station numbers are in inches 20.60 3. Acoustic blanket thickness is 38.1 mm (1.5 in.) in nose, 76.2 mm (3.0 in.) on large cylinder, and 38.1 mm (1.5 in.) on small cylinder 20 deg 4. MDA requires definition of spacecraft features within 50.8 mm (2.0 in) of payload envelope

5. Projections of spacecraft appendages below the spacecraft separation plane may be permitted but must be coordinated with the Delta Program Office

733 15 deg 28.84

6487 255.51

2896 Dia 114.00

1953 8488 76.91 334.17

2540 Dia 100.00

308 12.12

2184 Dia 86.00 1553 61.13

Sta 485.21 Separation Plane for 6915 PAF

Sta 500.21

Sta 553.39

2438 dia 96.00

Figure 3-4. Spacecraft Envelope, 2.9-m (9.5-ft) Diameter, Two-Stage Configuration (6915 PAF) (Download Figure) 3-6 02745REU7

CL of Air Conditioning Door I mm 43° 23' in

II IV

ContaminationÐFree Separation Joint

III View A-A 293 R 11.54

AA Sta 203.39

Sta 324.90

A/C Inlet Door Sta 356.90 8890 350

3693 145.38

24-in-dia Spacecraft Access Door (as required)

Sta 470.28 18-in-dia Access Door 2 Places Sta 506.30

9.75°

Sta 553.39 Inside Skin Dimensions

Figure 3-5. Profile, 3-m (10-ft) Composite Fairing (Download Figure) 3-7 smooth inside skin allows the flexibility to install three- and two-stage configurations. These figures mission-unique access doors almost anywhere in reflect envelopes for existing attach fittings and the cylindrical portion of the fairing. An RF win- assume that the spacecraft stiffness recommended dow can be accommodated, similar to mission- in Section 4.2.3.2 is maintained. Use of the portion unique access doors. All these requirements are to of the envelope shown in Figure 3-6 that is below be coordinated with the Delta Program office. the separation plane, and local protuberances out- The bisectors are joined by a contamination-free side the envelopes presented, require coordination linear piston/cylinder thrusting separation system and approval of the Delta Program Office. that runs longitudinally the full length of the fairing. 3.4 THE 3-m (10L-ft) DIAMETER Two functionally redundant explosive bolt assem- SPACECRAFT FAIRING blies provide the structural continuity at the base The 3-m (10-ft) diameter 10L fairing is available ring of the fairing. for spacecraft requiring a longer envelope than the The fairing bisectors are jettisoned by actuation standard 3-m (10-ft) fairing described in of the base separation nuts, and by the detonating Section 3.3. This fairing (Figure 3-8) is also a fuse in the thrusting joint cylinder rail cavity. A composite sandwich structure which separates into bellows assembly within each cylinder rail retains bisectors. The cylindrical section is 0.9 m (3 ft) the detonating-fuse gases to prevent contamination longer than its counterpart discussed in Section 3.3. of the spacecraft during the fairing separation event. The overall length is 0.4 m (1.3 ft) longer than the Two standard 18-in.-diameter access doors are standard 3-m (10-ft) fairing. part of the baseline fairing configuration for sec- Other than the difference in length, the discus- ond-stage access (Figure 3-5). To satisfy spacecraft sion in Section 3.3 also applies to the extended- requirements, additional standard 24-in.-diameter length 10L fairing. removable doors can be provided in the fairing The allowable static spacecraft envelopes within cylindrical section to permit access to the spacecraft the fairing are shown in Figures 3-9 and 3-10 for the following fairing installation. The quantity and three- and two-stage configurations. These figures location of additional access doors must be coordi- reflect the envelopes for existing attach fittings, and nated with the Delta Program Office. it is assumed that the spacecraft stiffness recom- Acoustic absorption blankets are provided on the mended in Section 4.2.3.2 is maintained. Use of the fairing interior. Typical blanket configurations are portion of the envelope that is below the separation described in Table 3-1. plane and local protruberances outside the enve- The allowable static spacecraft envelopes within lopes require coordination and approval of the the fairing are shown in Figures 3-6 and 3-7 for the Delta Program Office.

3-8 03058REU7

Fairing Envelope Sta 203.39 Sta 208.37 Usable Payload Envelope 305 Sta 215.99 R 12.0

Usable Envelope Below Separation Plane

Attach Fitting

Motor

Acoustic Blankets 5028 mm 6176 All dimensions are in 197.96 in. R 243.14

All station numbers are in inches Acoustic blanket thickness Sta 305.97 per Table 3-1

MDA requires definition of spacecraft features within 50.8 mm (2.00 in.) of payload envelope 2743 Projections of spacecraft dia 108.0 appendages below the 2743 spacecraft separation plane 107.98 Sta 373.45 may be permitted, but must be

coordinated with the Delta Program Office 3056 dia 120.3

Sta 413.95

Spacecraft Separation Plane for 3712 PAF 2743

108.0 dia

Sta 500.21 4563

179.64 2743 dia 108.0

1243

dia Sta 553.39

48.93 2438 940 dia dia 96.0 37.00 102 Inside Skin Dimensions 724 dia 4.00 28.50

Sta 413.95

533 21.0 15 deg 647 R 25.49 73 2.88 15 deg

Figure 3-6. Spacecraft Envelope, 3-m (10 ft) Diameter Fairing, Three-Stage Configuration (3712 PAF) (Download Figure) 3-9 02742REU7

Sta 203.39 305 R Sta 208.37 12.0

Sta 215.99 6838 269.22

6176 R 243.14

Sta 305.97

4553 Sta 373.45 2743 179.24 dia

108.0 3056 dia

120.3 Sta 485.21 Spacecraft Separation

Plane for 6915 PAF

Sta 500.21 Second Stage 4563

Interface Plane 179.64

Sta 553.39 2438 dia 96.0 Inside Skin Dimensions

Notes:

Fairing Envelope mm 1. All dimensions are in Usable Payload in. Envelope 2. All station numbers are in inches Attach Fitting 3. Acoustic blanket thickness per Table 3-1

Acoustic Blankets 4. MDA requires definition of spacecraft features within 50.8 mm (2.00 in.) of payload envelope Projections of spacecraft appendages below the spacecraft separation plane may be permitted but must be coordinated with the Delta Program Office

Figure 3-7. Spacecraft Envelope, 3-m (10-ft ) Diameter Fairing, Two-Stage Configuration (6915 PAF) (Download Figure) 3-10 02751REU7

C L of Air Conditioning Door I mm ° 43 23' in

II IV

Contamination-Free Separation Joint

III View A-A 285 R 11.21 3-m-dia (10-ft) Fairing 10L (Under Development)

AA Sta 189.12

Nose Cone

Sta 286.34

9252 A/C Inlet Door Sta 356.90 364.27

3-m-dia (10-ft) Cylinder 4672 183.94

24-in-dia Spacecraft Access Door (as required)

Sta 470.28 18-in-dia Access Door (2 Places) Sta 506.30

9.75°

Sta 553.39 2.4-m-dia (8-ft) Base Inside Skin Dimensions

Figure 3-8. Profile, 3 m (10-ft) 10L Composite Fairing (Download Figure) 3-11 02802REU7 Sta 189.12 Fairing Envelope 305 R Usable Payload 12.0 Sta 201.12 Envelope Usable Envelope Below Separation Plane

Attach Fitting Motor 32.1 deg Acoustic Blankets 32.1° mm All dimensions are in in. Sta 275.34

All station numbers are in inches 9252

Acoustic blanket thickness 364.27 per Table 3-1 5711 2743 dia 224.83 MDA requires definition of 108.0 spacecraft features within 50.8 mm 5406 (2.00 in.) of payload envelope 212.83 Projections of spacecraft 3660 144.08 appendages below the 3056 spacecraft separation plane dia 120.3 may be permitted, but must be coordinated with the Delta

Program Office

Sta 413.95 Spacecraft Separation Plane

for 3712 PAF 2743

108.0 dia

2743 dia 108.0 Fairing Separation Plane Sta 553.39 1243 dia

48.93 2438 940 dia dia 96.0 37.00 Inside Skin Dimensions 102 724 dia 4.00 28.50

Sta 413.95

533 21.0 15 deg 647 R 25.49 73 2.88 15 deg

Figure 3-9. Spacecraft Envelope, 3-m (10-ft) Diameter 10L Fairing, Three-Stage Configuration (3712) PAF (Download Figure) 3-12 03057REU7 305 R 12.0 Sta 189.12 Sta 194.72

Sta 201.12

32.1 deg

Sta 275.34

9252 364.27

7216 284.09

2743 dia 108.0 5331

209.87 3056 dia

120.3

Sta 485.21 Spacecraft Separation

Plane for 6915 PAF

Sta 500.21 381

15.00

Sta 553.39 2438 dia 96.0

Inside Skin Dimensions

Notes:

Fairing Envelope mm 1. All dimensions are in Usable Payload in. Envelope 2. All station numbers are in inches Attach Fitting 3. Acoustic blanket thickness per Table 3-1 Acoustic Blankets 4. MDA requires definition of spacecraft features within 50.8 mm (2.00 in.) of payload envelope Projections of spacecraft appendages below the spacecraft separation plane may be permitted but must be coordinated with the Delta Program Office

Figure 3-10. Spacecraft Envelope, 3-m (10-ft) Diameter 10L Fairing, Two-Stage Configuration (6915) PAF (Download Figure) 3-13 Section 4 H54869 SPACECRAFT ENVIRONMENTS

Lanyard Disconnects This section describes the launch vehicle envi- Fairing Wall (Inside) ronments to which the spacecraft is exposed during prelaunch activities and launch. Section 4.1 dis- Air Conditioning cusses prelaunch environments for processing Air Flow Duct facilities at both eastern and western ranges. Sec- tion 4.2 presents the Delta II launch and flight Air Conditioning Inlet Diffuser environments for the spacecraft. Air Flow 4.1 PRELAUNCH ENVIRONMENTS

4.1.1 Spacecraft Air Conditioning The environment that the spacecraft experiences

during its processing is carefully controlled for Air conditioning duct and diffuser temperature, relative humidity, and cleanliness. This system is ejected at liftoff includes the processing conducted before the space- Figure 4-1. Payload Air Distribution System craft is mounted on the Delta II, while it is in the mobile service tower (MST) white room, and after systems for fairing air conditioning. At SLC-2, the it is encapsulated within the payload fairing. fairing air conditioning system is redundant. At Air conditioning is supplied to the spacecraft via VAFB, emergency standby power can be arranged an umbilical after the spacecraft and fairing are if requested by the spacecraft agency. A GN purge mated to the Delta II. If an environmental shroud is 2 line to the spacecraft can be accommodated through required around the spacecraft prior to fairing in- the air-conditioning duct after fairing installation. stallation, it receives the same fairing air. The The air-conditioning duct is in the Quad I half of the spacecraft air-distribution system (Figure 4-1) pro- fairing. Unique mission requirements or equipment vides air at the required temperature, relative hu- should be coordinated with the Delta Program midity, and flow rate. The spacecraft air-distribution Office. system utilizes a diffuser on the inlet air-conditioning Various payload processing facilities are avail- duct at the fairing interface. If required, a deflector able at the launch site for use by the payload agency. can be installed on this inlet to direct the airflow Environmental control specifications for these away from sensitive spacecraft components. The facilities are listed in Tables 4-1 and 4-2 for eastern air-conditioning umbilical is pulled away at liftoff and western ranges, respectively. The facilities used by lanyard disconnects, and the access door on the would depend on spacecraft program requirements. fairing automatically closes. The air is supplied to the payload at a maximum flow rate of 1500 scfm. 4.1.2 MST White Room It flows downward around the spacecraft and is The white room is an environmentally controlled discharged below the second stage through vents in room located at the upper levels within the MST the interstage. At SLC-17, both pads have backup which has provisions for maintaining spacecraft 4-1 Md Table 4-1. Eastern Range Facility Environments Location Temperature Relative humidity Filtration Building AE Cleanroom complex 22.2 ± 1.76°C (72 ± 3°F) 55 ± 5% Class 10,000(1) Handling cans Mobile Note(2) Not controlled(3) Not controlled(3) MST(5) SLC-17A and SLC-17B 23.9 ± 2.8°C (75 ± 5°F) 45 ± 5% 99.97% of all particles white room (all doors over 0.3 µm closed) Class 100,000(5) Fairing and environmental Any specified between 50% max(6) 99.97% of all particles shroud(7) 15.5 and 26.6 ± 2.8°C over 0.3 µm (60 and 80 ± 5°F) Class 100,000(5) Astrotech Buildings 1 and 2: airlock, 23.9± 2.8°C (75 ± 5°F) 50 ± 5% Class 100,000(5) high bays Storage bays 21°C to 25.5°C 55% max Commercial standard (70°F to 78°F) SAEF 2 Airlock 21.7 ± 3.3°C (71 ± 6°F) 50% max Class 100,000(5) High bay 21.7 ± 3.3°C (71 ± 6°F) 50% max Class 100,000(5) Low bays 21.7 ± 3.3°C (71 ± 6°F) 50% max Class 100,000(5) Test cells 21.7 ± 3.3°C (71 ± 6°F) 50% max Class 100,000(5) Cargo hazardous Airlock 21 ± 2.8°C (70 ± 5°F) 30 to 50% Class 100,000(5) processing facility Hazardous operations bay 21 ± 2.8°C (70 ± 5°F) 30 to 50% Class 100,000(5) Note: The facilities listed can only lower the outside humidity level. The facilities do not have the capability to raise outside humidity levels. These numbers are provided for planning purposes only. Specific values should be obtained from controlling agency. (1)Class 1,000 obtainable with restrictions. (2)Passive temperature control provided by operational constraints. (3)Dry gaseous nitrogen purge per MIL-P-2740IC, Type 1, Grade A. (4)A backup system exists for the MST white room air conditioning. (5)FED-STD-209D. (6)50% relative humidity can be maintained at a temperature of 62°F (16.7°C). At higher temperatures, the relative humidity can be reduced by drying the conditioned air to a minimum specific humidity of 40 grains of moisture per 0.45 kg (1 lb) of dry air. (7)Fairing air customer temperature requirements over 22.8°C (73°F) should be coordinated with the Delta Program Office. M029, T007, 4/12/96, 11:41 AM

Table 4-2. Western Range Facility Environments Location Temperature Relative humidity Filtration Building 836(1) Spacecraft Lab 1 15.5 to 26.6 ± 1.2°C TBD Class 100,000(2) (60 to 80° ± 2°F) Spacecraft Lab 2 15.5 to 26.6 ± 1.2°C TBD Class 100,000(6) (60 to 80° ± 2°F) Hi-bay Note(1) TBD N/A Building 1610 Hazardous Processing 17.7 to 32.2° ± 2.8°CC 40% to 50% ± 5% Class 100,000 Facility (65 to 80° ± 5°F) Handling can Mobile Note(3) Not controlled (4) Not controlled (4) California Space Port Payload Checkout Cells 23.9 ± 2.8°C(75 ± 5°) 30Ð50% Class 100,000(6) Astrotech Payload Processing 15.5 to 26.6 ± 1.2°C(5) 35-60% ± 10(5) Class 100,000(6) Rooms (60 to 80° ± 2°F) functional 10,000 MST SLC-2 white room (all 21.1 ± 2.8°C (70 ± 5°F) 30% to 50% 99.97% of all particles doors closed) over 0.3 µm Class 100,000 Fairing and Any specified between 15.5 30% to 50% 99.97% of all particles (7) ± ° environmental shroud and 23.9 2.8 C (60 and over 0.3 µm ± ° 75 5 F) Class 100,0000(6) (1)These numbers are provided for preliminary planning purposes only. Payload users should contact NASA for specific capabilities of listed NASA facilities (2)Class 10,000 with strict controls (3)Passive temperature control provided by operational constraints (4)Dry and gaseous nitrogen purge per MIL-P-27401C, Type 1, Grade A (5)Controlled to customer requirement within range (6)Fed-Std-209D (7)Fairing air customer temperature requirements over 22.8°C (73°F) should be coordinated with the Delta Program Office

4-2 Md cleanliness. White room environments are listed in craft status may also be obtained from the vehicle Table 4-1 for pads A and B at SLC-17 and in Table telemetry system. 4-2 for SLC-2 at VAFB. SLC-17 also has a person- At the eastern and western ranges, the electro- nel changeout room for each pad adjacent to the magnetic environment to which the satellite is ex- white room. posed results from the operation of range radars and

4.1.3 Radiation and Electromagnetic the launch vehicle transmitters and antennas. The Environments maximum RF environment at the launch site is The Delta II transmits on several frequencies to controlled through coordination with the range and provide launch vehicle telemetry and beacon signals with protective masking of radars. The launch pads to the appropriate range tracking stations. It also has are protected to an environment of 10 V/m at uplink capability to onboard command receiver frequencies from 14 kHz to 40 GHz and 20 V/M in decoders (CRDs) for command destruct capability. the C-band frequency of the range tracking radars. Two S-band telemetry systems (one each on the An RF hazard analysis is performed to ensure second and third stage), two CRD systems on the that the satellite transmitters are compatible with second stage, and a C-band transponder (beacon) the vehicle avionics and ordnance systems. An RF on the second stage are provided. The radiation characteristics of these systems are shown in Table compatibility analysis is also performed to verify 4-3. The RF systems are switched on prior to launch that the vehicle and satellite transmitter frequencies and remain on until stage separation and battery do not have interfering intermodulation products or depletion. Spacecraft launch environment data such image rejection problems. as low- and high-frequency vibration, acceleration Payload agencies should contact the Delta Pro- transients, shock velocity increments, and space- gram Office for induced RF environments.

Table 4-3. Delta II Transmitter Characteristics Second-stage T/M radiation Third-stage T/M radiation Second-stage C-band characteristics characteristics beacon characteristics ■ Transmitter ● Nominal frequency 2241.5 MHz 2252.5 MHz 5765 MHz (transmit) 5690 MHz (receive)

● Power output 2.0 W min 5.0 W min 400 W min

● Modulation bandwidth ±160 kHz at 20 dB ±70 kHz at 20 dB 6 MHz at 6 dB ±650 kHz at 60 dB ±250 kHz at 60 dB

● Stability +67 kHz max +67 kHz max 3 MHz max ■ Antenna ● Type Cavity-backed slot Circumferential belt Transverse slot, dipole loaded

● Polarization Essentially linear parallel to Essentially linear parallel to Left hand circular booster roll axis booster roll axis ■ Locationˇ 316 deg (looking aft) Belt at Sta 438 153 deg (looking aft) Ð Sta 559 Ð Sta 559 143 deg (looking aft) Ð Sta 559 ● Pattern Nearly omnidirectional Nearly omnidirectional Nearly omnidirectional

● Gain +2.35 dB max +3 dB max +6 dB max M029, T009, 7/31/95, 2:14 PM 4-3 Md 4.1.4 Electrostatic Potential C. The fairing is cleaned using alcohol and then During ground processing, the spacecraft must inspected for cleanliness prior to spacecraft encap- be equipped with an accessible ground attachment sulation. Table 4-4 provides MDA STP0407 visible cleanliness (VC) levels with their NASA SN- point to which a conventional alligator-clip ground C-0005 equivalency. It defines the cleanliness strap can be attached. Preferably, the ground attach- levels available to payload customers. The standard ment point is located on or near the base of the level for a mission is VC 2. Other cleanliness levels spacecraft, at least 31.8 mm (1.25 in.) above the must be negotiated with the Delta Program Off separation plane. The vehicle/spacecraft interface provides the conductive path for grounding the spacecraft to the launch vehicle. Therefore, dielec- Table 4-4. Cleanliness level Definitions MDA STP0407-0X NASA SN-C-0005 tric coating should not be applied to the spacecraft VC 1 None VC 2 VC Standard interface. The electrical resistance of the spacecraft VC 3 VC Highly Sensitive to PAF interface surfaces must be 0.0025 ohm or VC 4 VC Sensitive + UV (Closest equivalent. MDA is more critical) less and is verified during spacecraft to PAF mate. VC 5 VC Highly Sensitive VC 6 VC Highly Sensitive + UV (Reference MIL-B-5087B, Class R.) M046 t1

4.1.5 Contamination and Cleanliness Cleanliness Level Definitions Cleanliness conditions provided for the Delta II VC 1 - All surfaces shall be visibly free of all payloads represent minimum cleanliness condi- particulate and nonparticulate visible to the normal tions available. The following guidelines and prac- unaided (except for corrected vision) eye. Particu- tices from prelaunch through spacecraft separation late is defined as matter of miniature size with provide the minimum Class 100,000 cleanliness observable length, width and thickness. conditions (per Federal Standard 209B): Nonparticuate is film matter without definite di- A. Precautions are taken during manufacture, mension. Inspection operations shall be performed assembly, test, and shipment to prevent contami- under normal shop lighting conditions at a maxi- nant accumulations in the Delta II upper-stage area, mum distance of 3 feet. fairing, and PAF. VC 2 - All surfaces shall be visibly free of all B. Encapsulation of the payload into the han- particulate and nonparticulate visible to the normal dling can is performed in a facility that is environ- unaided (except for corrected vision) eye. Particu- mentally controlled to Class 100,000 conditions. late is identified as matter of miniature size with All handling equipment is cleanroom compatible observable length, width and thickness. and is cleaned and inspected before it enters the Nonparticulate is a film matter without definite facility. These environmentally controlled condi- dimension. Inspection operations shall be per- tions are available for all remote encapsulation formed at incident light levels of 50 foot candles facilities and include SLC-17 and SLC-2. A han- and observation distances of 5 to 10 feet. dling can is used to transport the payload to the VC 3 - All surfaces shall be visibly free of all white room and provides environmental protection particulate and nonparticulate visible to the normal for the payload. unaided (except for corrected vision) eye. Particu- 4-4 Md late is identified as matter of miniature size with with UV lamps. Cleaning must be done in a Class observable length, width and thickness. 100,000 clean room or better. Nonparticulate is a film matter without definite D. Personnel and operational controls are em- dimension. Incident light levels of 100 to 200 foot ployed during spacecraft encapsulation to maintain candles at observation distance of 18 inches or less. spacecraft cleanliness. VC 4 - All surfaces shall be visibly free of all E. The payload agency may provide a protective particulate and nonparticulate visible to the normal barrier (bag) around the spacecraft prior to encapsu- unaided (except for corrected vision) eye. Particu- lation in the handling can. late is identified as matter of miniature size with F. A contamination barrier (bag) is installed observable length, width and thickness. around the handling can immediately following Nonparticulate is a film matter without definite encapsulation operations. An outer bag is installed dimension. This level requires no particle count. for transportation. A nitrogen purge is provided to the handling can during transport. The source of incident light shall be a 300 watt drop G. A payload environmental shroud can be pro- light (explosion-proof) held at distance of 5 feet, vided in the white room for the spacecraft prior to maximum, from the local area of inspection. There fairing installation. This shroud enables the space- shall be no hydrocarbon contamination on surfaces craft to be showered with class 10,000 fairing air. specifying VC 4 cleanliness.

VC 5 - All surfaces shall be visibly free of all 4.2 LAUNCH AND FLIGHT ENVIRONMENTS particulate and nonparticulate visible to the normal unaided (except for corrected vision) eye. Particu- 4.2.1 Fairing Internal Pressure Environment late is identified as matter of miniature size with As the Delta vehicle ascends through the atmo- observable length, width and thickness. sphere, the fairing is vented through a 64.5 cm2 (10 Nonparticulate is a film matter without definite in.2) opening in the interstage and other leak paths dimension. Incident light levels of 100 to 200 foot in the vehicle. The expected extremes of internal candles at observation distance of 6 to 18 inches. pressure during ascent are presented in Figure 4-2 Cleaning must be done in a Class 100,000 clean for the 2.9-m (9.5-ft) and 3-m (10-ft) fairings. room or better. VC 6 - All surfaces shall be visibly free of all 4.2.2 Thermal Environment particulate and nonparticulate visible to the normal Prior to and during launch, the Delta II payload unaided (except for corrected vision) eye. Particu- fairing and upper stages contribute to the thermal late is identified as matter of miniature size with environment of the spacecraft. observable length, width and thickness. 4.2.2.1 Payload Fairing Thermal Environment. Nonparticulate is a film matter without definite Upon PLF installation, air conditioning is provided dimension. Incident light levels of 100 to 200 foot at temperatures ranging from 12.8°C to 26.7°C candles at observation distance of 6 to 18 inches. (55°F to 80°F) depending on mission requirements. Additional incident light requirements are 8 watts Variations in this range can be accommodated and minimum of long wave ultraviolet light at 6 to 18 should be coordinated with the Delta Program Of- inches observation distance in a darkened work fice. Details of the air conditioning capability of the area. Protective eyeware may be used as required launch sites are provided in Section 4.1. 4-5 Md H54607.2 16

Minimum Pressure Maximum Pressure 10

6 Compartment Pressure (psi) Pressure Compartment

0 0 10203040506070 Time (sec)

Figure 4-2. Delta II Payload Fairing Compartment Absolute Pressure Envelope

The ascent thermal environments of the Delta II keted. There may be slight variations in blanket fairing surfaces facing the spacecraft are shown in coverage areas based on mission unique require- Figures 4-3 and 4-4. Typical maximum tempera- ments. Inclusion of an RF window in the 9.5-ft PLF ture histories of the fairing skin and the inboard conical section results in a local increase in acoustic facing surface of the acoustic blanket are shown for blanket temperature inboard of the RF window as the 2.9-m (9.5-ft) , and 3-m (10-ft) fairing configu- shown in Figure 4-3. rations on the 792X vehicle. Temperatures are The fairing skin temperature is representative of provided for both the PLF conical section and the the radiation environment to the spacecraft in unblanketed areas such as the air conditioning inlet cylindrical section facing the spacecraft. PLF in- door, unblanketed access doors and blanket cutout board facing surface emissivity values are also provided. All temperature histories presented are regions. Maximum skin temperatures are shown in based on depressed versions of the trajectory Figures 4-3 and 4-4. (worst case). The 9.5-ft fairing frame temperatures are some- The acoustic blankets provide a relatively cool what less severe than skin temperatures. Informa- radiation environment by effectively shielding the tion regarding frame locations, exposure, and tem- spacecraft from the ascent heating in blanketed perature history is available on request. areas. Figures 4-3 and 4-4 depict the areas of the Unless otherwise requested, jettison of the fair- various Delta II fairings which are typically blan- ing will occur shortly after the theoretical free 4-6 Md H54473.1 AA Sparesyl Insulation on Nose Cap and Cone

38.1 mm/1.5 in. ThickAA Cone Acoustic Blanket 1 Aluminum Sector/ AA 1 Fiberglass Sector RF Window* AA (Aluminum Foil Removed From AA Fiberglass Cone) 76.2 mm/3.0 in. Thick 9.5-ft. Cyclinder Acoustic BlanketAA AA 38.1 mm/1.5 in. Thick Internal Surface Emittance Acoustic BlanketAA Nose cap, aluminum cone, AA fiberglass cone with aluminumfoil 0.10 Sparesyl Insulation RF window (fiberglass cone without foil) 0.90 on Separation Rail 8-ft. Cyclinder 9.5-ft cylinder 0.10 8-ft cylinder 0.25 *Size and location vary with spacecraft Acoustic blanket 0.86

600 300

9.5-ft Cylinder Skin 500 250

400 200

8-ft Aft Cylinder Skin

150 300

Cone Skin

Temperature (¡C) Temperature Temperature (¡F) Temperature

Blanket at RF Window 100 200 Cone Blanket

50 100

Cylinder Blanket

Spacecraft at 70¡F with emittance of 0.1 0

0 0 50 100 150 200 250 300 Time From Liftoff (sec)

Figure 4-3. Predicted Maximum Internal Wall Temperatures and Internal Surface Emittance (9.5-ft Fairing) 4-7 Md H55170.5

Acoustic Blanket Thickness Sta 208.37 AAAAA Sparesyl Insulation on Nose Cap and Cone AAAAA (Skin and Separation Rail) AAAAA AAAAA AAAAA AAAAA 76.2 mm/3.0 in. AAAAA Sparesyl Insulation on AAAAA Separation Rail

AAAAA Internal Surface Emittance AAAAA Nose cap, cone, unblanketed skin 0.90 Acoustic blanket 0.10

AAAAA Unblanketed rail 0.10 AAAAA Sta 553.39 AAAAA

160 A A A A A

A A SkinA Inside SurfaceA A 140 and Separation AAAAAAAAAAAAAA A A RailA A

120AAAAAAAAAAAAAA A A A A A A A A A 100AAAAAAAAAAAAAA A A A A Spacecraft at 70¡F with emmitance of 0.1

Temperature (¡F) Temperature AAAAAAAAAAAAA80 A A A A A A A A A A 60 AAAAAAAAAAAAAA A A ABlanket A A A A A A 40 0 50 100 150 200 250 300 Time (sec)

Figure 4-4 Predicted Maximum Internal Wall Temperatures and Internal Surface Emittances (10-ft Fairing) 4-8 Md molecular heating for a flat plate normal to the free energy from the Delta third stage spin rockets stream drops below 0.1 Btu/ft2-sec (1135 W/m2) during burn and from the third stage motor during based on the 1962 US Standard Atmosphere. and after burn. Because of the high temperature of the exhaust gases, the radiation from the plumes 4.2.2.2 On-Orbit Thermal Environment. During may be significant if there are sensitive compo- coast periods, the Delta II can be oriented to meet nents located at the spacecraft base. specific sun angle requirements. A slow roll during The third-stage spin rocket plumes subject the a long coast period can also be used to moderate spacecraft to a maximum heat flux of 2840 W/m2 orbital heating and cooling. The Delta II roll rate for (0.25 Btu/ft2-sec) at the spacecraft/third stage sepa- thermal control is typically between 1 and 3 ration plane. This heat flux is a pulse of 1-sec deg/sec. duration. The third-stage motor plume subjects the space- 4.2.2.3 Spacecraft/Launch Vehicle Interface. 2 The spacecraft is required to provide interface tem- craft to maximum heat flux of 2044 W/m (0.18 Btu/ft2-sec). The motor burns for approximately peratures for the injection period assuming an adia- 86 sec. Plume maximum heat flux is plotted versus batic interface. MDA will provide launch vehicle radial distance in Figure 4-5. The variation of the interface temperatures also assuming an adiabatic heat flux with time during the third-stage burn is interface. shown in Figure 4-6. 4.2.2.4 Upper Stage Induced Thermal Envi- After third-stage motor burnout, the titanium ronments. The spacecraft receives radiant heat motor case temperature rises rapidly, as shown in

H54472.1

0.20

0.18 L 2000

0.16

0.14

1500 2

sec

Btu

2 ( ) (

ft 0.12

( ) (

max

Q max

Q 0.10

1000 0.08

0.06

0.04 500 25 30 35 40 45 50 55 60

Distance From Centerline, L (in.) Figure 4-5. Predicted Star 48B Plume Radiation at the Spacecraft Separation Plane Versus Radial Distance 4-9 Md H54470.1

1.0

0.9

0.8

max Q/Q 0.7

0.6

0.5 0 10 20 30 40 50 60 70 80 90

Burn Time (sec) Figure 4-6. Predicted Star 48B Plume Radiation at the Spacecraft Separation Plane Versus Burn Time

Figure 4-7. By 200 sec after motor ignition (typical plane. Any spacecraft instruments which protrude spacecraft separation time), the case temperature below the interface plane should be evaluated for reaches 299°C (570°F). Case temperature eventu- proximity to the NCS thruster. Information regard- ally reaches a peak of 343°C (650°F) at approxi- ing this plume can be provided as necessary. mately 265 sec after motor ignition. The temperature history shown is the maximum expected along the forward portion of the motor case and corre- 4.2.3 Flight Dynamic Environment sponds to a 7925 Delta II class spacecraft weight The acoustic, sinusoidal, and shock environ- greater than 820 kg (1800 lbs), which yields ap- ments provided in Sections 4.2.3.3 through 4.2.3.5 proximately 10 pounds of slag. For lighter space- are based on maximum flight levels for a 95th craft weights [less than 820 kg (1800 lbs)], the case percentile statistical estimate. temperature peaks at approximately 427°C (800°F) and produces approximately 16 pounds of slag due 4.2.3.1 Steady-State Acceleration. For the to the higher acceleration. Mission users should Delta 7320 and 7920 vehicles, the maximum axial contact the Delta Program Office for greater detail. acceleration occurs at the end of the first-stage burn The external surface emissivity of the titanium (MECO). For a three-stage Delta vehicle, the motor case is 0.34. maximum steady-state acceleration occurs at the The hydrazine thruster plume of the third-stage end of third-stage flight for payloads up to 885 kg nutation control system (NCS) does not introduce (1950 lb). Above this weight the maximum accel- any significant heating to the spacecraft interface eration occurs at MECO. A plot of steady-state 4-10 Md H54471.3

1000 500

450 800 Light Spacecraft 400

350

600 Heavy Spacecraft 300

250

Forward Dome Temperature

400 200 Temperature (¡C) Temperature Temperature (¡F) Temperature From A to B

66 deg B 150

200 100 A ε = 0.34 50

100 0 100 200 300 Time From Third-Stage Ignition (sec) Figure 4-7. Predicted Star 48B Motor Case Soakback Temperature axial acceleration at MECO versus payload weight stage spacecraft) in the lateral axis for a spacecraft is shown in Figure 4-8 and is respresentative for the hard-mounted at the spacecraft separation plane acceleration at MECO for the 2.9- and 3-m (9.5 and (without PAF and separation clamp). In addition, 10 ft) fairings. Steady-state axial acceleration ver- secondary structure mode frequencies should be sus spacecraft weight at third-stage motor burnout above 35 Hz to avoid undesirable coupling with is shown in Figure 4-9. launch vehicle modes and/or large fairing-to-space- 4.2.3.2 Combined Loads. Dynamic excita- craft relative dynamic deflections. The spacecraft tions, which occur predominantly during liftoff and design limit load factors presented in Table 4-5 are transonic periods of Delta flight, are superimposed applicable for spacecraft meeting the above funda- on steady-state accelerations to produce combined mental frequency criteria. For very flexible space- accelerations that must be used in the spacecraft craft, the combined accelerations and subsequent structural design. The combined spacecraft accel- design limit-load factors could be higher than shown, erations are a function of launch vehicle character- and the user should consult the Delta Program istics as well as spacecraft dynamic characteristics Office so that appropriate analyses can be per- and mass properties. To avoid dynamic coupling formed to better define loading conditions. between low-frequency vehicle and spacecraft 4.2.3.3 Acoustic Environment. The maximum modes, the stiffness of the spacecraft structure acoustic environment for the spacecraft occurs dur- should produce fundamental frequencies above 35 ing liftoff and transonic flight, and the duration of Hz in the thrust axis and 15 Hz (12 Hz for a two- the maximum environment is less than 10 seconds. 4-11 Md H54870.1 8.0AAAAAAAAAAAAAA Note: Second stage payload weight is defined as the sum of the AAAAAAAAAAAAAANote: weights of the spacecraft, PAF, third stage, and spin table. Note: The PAF, fully loaded third-stage motor, and spin table 7.5AAAAAAAAAAAAAANote: weight is 2308.8 kg (5,090 lb.) AAAAAAAAAAAAAA 7.0AAAAAAAAAAAAAA 3-Sigma High

AAAAAAAAAAAAAANominal 6.5AAAAAAAAAAAAAA

6.0AAAAAAAAAAAAAA Steady-State Acceleration (g’s) Acceleration Steady-State AAAAAAAAAAAAAA

5.5

5.0 02000 4000 6000 8000 10000 12000 Weight of Second-Stage Payload (lb)

01000 2000 3000 4000 5000 Mass of Second-Stage Payload (kg) Figure 4-8. Axial Steady-State Acceleration at MECO Versus Second-Stage Payload Weight

H54871.1 18

16AAAAA

14AAAAA

12AAAAA AAAAA 10AAAAA 8 3-Sigma High AAAAANominalAAAAAAA

6 AAAAAAA Steady-State Acceleration (g’s) Acceleration Steady-State 4 AAAAAAAAAAAA 2

0 500 1000 1500 2000 2500 3000 3500 4000 4500 Spacecraft Weight (lb)

200400 600 800 1000 1200 1400 1600 1800 2000 Spacecraft Mass (kg)

Figure 4-9. Axial Steady-State Acceleration at Third-Stage Burnout 4-12 Md Table 4-5. Spacecraft CG Limit-Load Factors (g) The spacecraft acoustic environment is a function Liftoff/Transonic of the configuration of the launch vehicle, the Axis Two-Stage Three-Stage MECO Lateral ±2.0 ±2.5 ±0.1 fairing, and the fairing acoustic blankets. Section 3 ±2.5(1) ±3.0(1) Thrust +2.8/-0.2(2) +2.8/-0.2(2) +6.2±0.6(3) defines the fairing blanket configurations. Table 4-6 (1)Lateral load factor to provide correct bending moment at spacecraft separation plane. identifies the figures that define the spacecraft (2)Plus indicates compression load and minus indicates acoustic environment for several versions of the tension load. (3)Axial load factor at MECO consists of a static Delta II. The maximum flight level spacecraft acous- component which is a function of spacecraft weight (Fig 4-8) and a dynamic component at a frequency of tic environments for blanketed region for the Delta ~17-18 Hz. The 6.2-g static value shown is based on a spacecraft weight of 1885 kg (4155 lb) for a 7900 series launch vehicle configurations are de- three-stage mission. M029, T006.2, 8/14/95, 2:14 PM fined in Figures 4-10 and 4-11 based on typical

Table 4-6. Spacecraft Acoustic Environment Figure Reference Delta II launch Fairing acoustic Spacecraft acoustic vehicle configuration Mission Fairing configuration blanket configuration environment type 7925 2-stage and 3-stage 9.5-ft diameter 3-in. configuration See Figure 4-10 7920 7325 7320 7925-10 2-stage and 3-stage 10-ft diameter 3-in. configuration See Figure 4-11 7920-10 7325-10 7320-10 M046, t4-5

H55384.4 150 Maximum Flight Levels (dB) One-Third Octave Center Frequency 3-Stage 2-Stage (Hz) Mission Mission 31.5 121.5 121.5 40 124 124 140 50 127 127 63 127.5 127.5 2 Stage 80 128.5 128.5 100 129 129.5 3 Stage 125 129.5 130.5 130 160 129.5 131 200 130 132 250 130 133 315 130 135 400 129 139 500 126.5 140.5 120 630 124 138 800 121 133

Sound Pressure Level (dB) Level Pressure Sound 1000 117 131 1250 114.5 130.5 1600 112 130.5 2000 109.5 128.5 110 2500 108 127 3150 106.5 127 4000 104.5 125 5000 104 124 6300 103 120.5 8000 102.5 119.5 100 31.5 63 125 250 500 1000 2000 4000 8000 10,000 102.5 118.5 One-Third Octave Band Center Frequency (Hz) OASPL 139.8 146.6 Duration 5 seconds 10 seconds

Figure 4-10. Predicted Delta II 7920 and 7925 9.5-ft Fairing Spacecraft Acoustic Environment 4-13 Md H55385.1 140 Maxunum Flight Levels (dB) One-Third Octave 3.0-in. Blanket Center Frequency Configuration (Hz) (dB) 31.5 119.5 40 122.5 130 50 126.5 63 128 2 Stage and 3 Stage 80 128.5 100 129.5 125 130 120 160 130 200 130.5 250 131.5 315 132.5 400 131.5 500 128 110 630 125 800 122 Sound Pressure Level (dB) Level Pressure Sound 1000 120 1250 118 1600 117 2000 116.5 100 2500 116 3150 115 4000 113.5 5000 111 6300 107 90 8000 103 31.5 63 125 250 500 1000 2000 4000 8000 10,000 100 One-Third Octave Band Center frequency (Hz) OASPL 140.8 Duration 5 seconds

Figure 4-11. Predicted Delta II 7920 and 7925 10-ft Fairing Spacecraft Acoustic Environment spacecraft with payload bay fills up to 60%. Launch spacecraft interface random vibration environment vehicles with payload bay fills above 80% will is not specified. experience approximately 1 1/2 dB higher levels. 4.2.3.4 Sinusoidal Vibration Environment. The The overall sound pressure level (OASPL) for each spacecraft will experience sinusoidal vibration in- acoustic environment is also shown in the figures. puts during flight as a result of Delta II launch and The maximum spacecraft acoustic environments ascent transients and oscillatory flight events. The for the Delta 7300 series launch vehicle configura- maximum flight level sinusoidal vibration inputs tions are 2.0 dB lower than those for the 7900 series are the same for all Delta II launch vehicle configu- vehicle because fewer GEM solid motors are ig- rations and are defined in Table 4-7 at the base of nited at liftoff. The acoustic environment produces the payload attach fitting. These sinusoidal vibra- the dominant high-frequency random vibration re- tion levels provide general envelope low-frequency sponses in the spacecraft, and a properly performed flight dynamic events such as liftoff transients, acoustic test is the best simulation of the acousti- transonic/maximum Q oscillations, pre-main en- cally-induced random vibration environment (See Table 4-7. Sinusoidal Vibration Levels Section 4.2.4.2). There are no significant high- Axis Frequency (Hz) Maximum flight levels frequency random vibration inputs at the PAF/ Thrust 5 to 6.2 1.27 cm (0.5 inch) double amplitude spacecraft interface that are generated by the Delta 6.2 to 100 1.0 g (zero to peak) Lateral 5 to 100 0.7 g (zero to peak) II launch vehicle; consequently, a Delta II PAF/ M046,T4-6 4-14 Md gine cutoff (MECO) sinusoidal oscillations, MECO The maximum flight level shock environments at transients, and second/third-stage events. the PAF/spacecraft interface defined in The sinusoidal vibration levels in Table 4-7 are Figures 4-12 through 4-15 are intended to aid in the not intended for use in the design of spacecraft design of spacecraft components and secondary primary structure. Limit load factors for spacecraft structure that may be sensitive to high-frequency primary structure design are specified in Table 4-5. pyrotechnic-shock. Typical of this type of shock, The sinusoidal vibration levels should be used in the level dissipates rapidly with distance and the conjunction with the results of the coupled dynamic number of joints between the shock source and the loads analysis (Table 8-3, Item 6) to aid in the component of interest. A properly performed sys- design of spacecraft secondary structure (e.g., solar tem-level shock test is the best simulation of the arrays, antennae, appendages, etc.) that may expe- high-frequency pyrotechnic shock environment rience dynamic loading due to coupling with Delta (see Section 4.2.4.4). II launch vehicle low-frequency dynamic oscilla- 4.2.4 Spacecraft Qualification and Accep- tions. Notching of the sinusoidal vibration input tance Testing. This section outlines a series of levels at spacecraft fundamental frequencies maybe environmental system-level qualification, accep- required during testing and should be based on the tance, and protoflight tests for spacecraft launched results of the vehicle coupled dynamic loads analy- on Delta II vehicles. The tests presented here are, by sis (see Section 4.2.4.3). necessity, generalized in order to encompass nu- 4.2.3.5 Shock Environment. The maximum shock merous spacecraft configurations. For this reason, environment at the PAF/spacecraft interface occurs each spacecraft project should critically evaluate its during spacecraft separation from the Delta II launch own specific requirements and develop detailed test vehicle and is a function of the PAF/spacecraft specifications tailored to its particular spacecraft. separation system configuration. Table 4-8 identi- Coordination with the Delta Program Office during fies the Figures that define the shock environment the development of spacecraft test specifications is at the spacecraft interface for various missions, encouraged to ensure the adequacy of the spacecraft PAF configurations, and types of separation sys- test approach (see Table 8-3, Item 5). tems. Shock levels at the PAF/spacecraft interface The qualification test levels presented in this due to other flight shock events such as stage section are intended to ensure that the spacecraft separation, fairing separation, and engine ignition/ possesses adequate design margin to withstand the shutdown are not significant compared to the space- maximum expected Delta II dynamic environmen- craft separation shock environment. tal loads, even with minor weight and design varia-

Table 4-8. Spacecraft Interface Shock Environment Figure Reference Mission PAF Spacecraft separation Spacecraft interface shock type configuration system type environment 3-Stage 3712A 37-in.-dia V-block clamp See Figure 4-12 3712B 3712C 2-Stage 6306 63-in. dia V-block clamp See Figure 4-13 2-Stage 6019 3 explosive separation nuts See Figure 4-14 2-Stage 6915 4 explosive separation nuts See Figure 4-14 2-Stage 5624 56-in.-dia V-block clamp See Figure 4-15 M046,T4-7 4-15 Md H55406.1 Shock Response Spectrum Q = 10 10,000

1500 Hz

4100 g

3000 Hz

1000

100 Peak Acceleration Response (g) Response Acceleration Peak

40 g

10 10 100 1000 10,000 Frequency (Hz)

Figure 4-12. Maximum Flight Spacecraft Interface Shock Environment-3712A, 3712B, 3712C Payload Attach Fitting

H55407.1 Shock Response Spectrum Q = 10 10,000

800 Hz

3000 g

3000 Hz

1000

100 Peak Acceleration Response (g) Response Acceleration Peak

10 10 100 1000 10,000 Frequency (Hz)

Figure 4-13. Maximum Flight Spacecraft Interface Shock Environment-6306 Payload Attach Fitting 4-16 Md H55408.2 Shock Response Spectrum Q = 10 10,000

5500 g

4000 Hz 5000 Hz

1700 Hz 2000 g

1000

100 Peak Acceleration Response (g) Response Acceleration Peak 350 Hz

10 10 100 1000 10,000 Frequency (Hz)

Figure 4-14. Maximum Flight Spacecraft Interface Shock Environment-6019 and 6915 Payload Attach Fitting H55409.1 Shock Response Spectrum Q = 10 10,000

900 Hz 3000 g

3000 Hz

1000

100 Peak Acceleration Response (g) Response Acceleration Peak

50 g

10 10 100 1000 10,000 Frequency (Hz)

Figure 4-15. Maximum Flight Spacecraft Interface Shock Environment-5624 Payload Attach Fitting 4-17 Md tions. The acceptance test levels are intended to After the test item is properly oriented and verify adequate spacecraft manufacture and work- mounted on the centrifuge, measurements and cal- manship by subjecting the flight spacecraft to maxi- culations must be made to assure the end of the test mum expected flight environments. The protoflight item nearest to the center of the centrifuge will be test approach is intended to combine verification of subjected to no less than 90 percent of the g-level adequate design margin and adequacy of spacecraft established for the test. If the g-level is found to be manufacture and workmanship by subjecting the less than 90 percent of the established g-level, the flight spacecraft to protoflight test levels, which are test item must be mounted further out on the centri- equal to qualification test levels with reduced dura- fuge arm and the rotational speed adjusted accord- tions. ingly or a larger centrifuge used so that the end of the test item nearest to the center of the centrifuge 4.2.4.1 Structural Load Testing. Structural load is subjected to at least 90 percent of the established testing is performed by the user to demonstrate the g-level. However, the opposite end of the test item design integrity of the primary structural elements (the end farthest from the center of the centrifuge) of the spacecraft. These loads are based on worst- should not be subjected to over 110 percent of the case conditions as defined in Sections 4.2.3.1 and established g-level. For large test items, exceptions 4.2.3.2. Maximum flight loads will be increased by should be made for load gradients based on the a factor of 1.25 to determine qualification test loads. existing availability of large centrifuges in com- A test PAF is required to provide proper load mercial or government test facilities. distribution at the spacecraft interface. The space- 4.2.4.2 Acoustic Testing. The maximum flight craft user shall consult the Delta Program Office level acoustic environments defined in before developing the structural load test plan and Section 4.2.3.3 are increased by 3.0 dB for space- shall obtain concurrence for the test load magnitude craft acoustic qualification and protoflight testing. in order to ensure that the PAF will not be stressed The acoustic test duration is 120 seconds for quali- beyond its load carrying capability. fication testing and 60 seconds for protoflight test- When the maximum axial load is controlled by ing. For spacecraft acoustic acceptance testing, the the third stage, radial accelerations due to spin must acoustic test levels are equal to the maximum flight be included. Spacecraft combined-loading qualifi- level acoustic environments defined in cation testing is accomplished by a static load test or Section 4.2.3.3. The acoustic acceptance test dura- on a centrifuge. Generally, static load tests can be tion is 60 seconds. The acoustic qualification, ac- readily performed on structures with easily defined ceptance, and protoflight test levels for the Delta I1 load paths, whereas for complex spacecraft assem- 7900 series launch vehicle configurations are de- blies, centrifuge testing may be the most economi- fined in Tables 4-9 through 4-11. As indicated, the cal. test levels for the Delta II 7300 series launch vehicle Test duration shall be 30 sec. Test tolerances and configurations are 2.0 dB lower than those for the mounting of the spacecraft for centrifuge testing is 7900 series launch vehicle. accomplished per Paragraph 4, Method 513, Mili- The acoustic test tolerances are +4 dB and -2 dB tary Standard 81OE, Environmental Test Methods, from 50 Hz to 2000 Hz. Above and below these dated 14 July 1989, which states: frequencies the acoustic test levels should be main- 4-18 Md Table 4-9. Acoustic Test Levels, Delta II 7925 Table 4-10. Acoustic Test Levels, Delta II 7920 9.5-ft Fairing, Three-Stage Mission, 9.5-ft Fairing, Two-Stage Mission, 3.0-in. Blanket Configuration 3.0-in. Blanket Configuration One-third One-third octave octave center Acceptance Qualification Protoflight center Acceptance Qualification Protoflight frequency test levels test levels test levels frequency test levels test levels test levels (Hz) (dB) (dB) (dB) (Hz) (dB) (dB) (dB) 31.5 121.5 124.5 124.5 31.5 121.5 124.5 124.5 40 124 127 127 40 124 127 127 50 126 129 129 50 127 130 130 63 127 130 130 63 129 132 132 80 128.5 131.5 131.5 80 130 133 133 100 129 132 132 100 130 133 133 125 129.5 132.5 132.5 125 130.5 133.5 133.5 160 129.5 132.5 132.5 160 131 134 134 200 130 133 133 200 132 135 135 250 130 133 133 250 133 136 136 315 130 133 133 315 135 138 138 400 129.5 132.5 132.5 400 139 142 142 500 127.5 130.5 130.5 500 140.5 143.5 143.5 630 125.5 128.5 128.5 630 138 141 141 800 124.5 127.5 127.5 800 133 136 136 1000 122 125 125 1000 131 134 134 1250 119 122 122 1250 130.5 133.5 133.5 1600 117.5 120.5 120.5 1600 130.5 133.5 133.5 2000 116.5 119.5 119.5 2000 128.5 131.5 131.5 2500 115.5 118.5 118.5 2500 127 130 130 3150 114 117 117 3150 127 130 130 4000 112.5 115.5 115.5 4000 125 125 125 5000 110.5 113.5 113.5 5000 124 127 127 6300 108.5 111.5 111.5 6300 120.5 123.5 123.5 8000 107 110 110 8000 119.5 122.5 122.5 10000 105.5 108.5 108.5 10000 118.5 121.5 121.5 OASPL 140.0 143.0 143.0 OASPL 146.7 149.7 149.7 Duration 60 sec 120 sec 60 sec Duration 60 sec 120 sec 60 sec Note: Delta II 7325 acoustic environment is 2.0 dB lower M046,T4-8 Note: Delta II 7320 acoustic environment is 2.0 dB lower M046,T4-9 tained as close to the nominal test levels as possible tions are defined in Tables 4-12 through 4-14 at the within the limitations of the test facility. The overall base of the payload attach fitting. sound pressure level (OASPL) should be main- The spacecraft sinusoidal vibration qualifica- tained within +3 dB and -1 dB of the nominal overall tion test consists of one sweep through the speci- test level. fied frequency range using a logarithmic sweep 4.2.4.3 Sinusoidal Vibration Testing. The maxi- rate of 2 octaves per minute. For spacecraft accep- mum flight level sinusoidal vibration environments tance and protoflight testing, the test consists of defined in Section 4.2.3.4 are increased by 3.0 dB (a one sweep through the specified frequency range factor of 1.4) for spacecraft qualification and using a logarithmic sweep rate of 4 octaves per protoflight testing. For spacecraft acceptance test- minute. The sinusoidal vibration test input levels ing, the sinusoidal vibration test levels are equal to should be maintained within + or - 10% of the the maximum flight level sinusoidal vibration envi- nominal test levels throughout the test frequency ronments defined in Section 4.2.3.4. The sinusoidal range. vibration qualification, acceptance, and protoflight When testing a spacecraft with a shaker in the test levels for all Delta II launch vehicle configura- laboratory, it is not within the current state of the 4-19 Md Table 4-11. Acoustic Test Levels, Delta II 7920 and art to duplicate the boundary conditions at the 7925 10-ft Fairing, Two- and Three-Stage Mission, 3.0-in. Blanket Configuration shaker input that actually occur in flight. This is One-third notably evident in the spacecraft lateral axis during octave center Acceptance Qualification Protoflight test, when the shaker applies large vibratory forces frequency test levels test levels test levels (Hz) (dB) (dB) (dB) to maintain a constant acceleration input level at 31.5 119.5 122.5 122.5 40 122.5 125.5 125.5 the spacecraft fundamental lateral test frequencies. 50 126.5 129.5 129.5 The response levels experienced by the spacecraft 63 128 131 131 80 128.5 131.5 131.5 at these fundamental frequencies during test are 100 129.5 132.5 132.5 usually much more severe than those experienced 125 130 133 133 160 130 133 133 in flight. The significant lateral loading to the 200 130.5 133.5 133.5 250 131.5 134.5 134.5 spacecraft during flight is usually governed by the 315 132.5 135.5 135.5 effects of spacecraft/launch vehicle dynamic cou- 400 131.5 134.5 134.5 500 128 131.5 131.5 pling. 630 125 128 128 800 122 125 125 Where it can be shown by a spacecraft launch 1000 120 123 123 vehicle coupled dynamic loads analysis that the 1250 118 121.5 121.5 1600 117 120 120 spacecraft or PAF/spacecraft assembly would ex- 2000 116.5 119.5 119.5 2500 116 119 119 perience unrealistic response levels during test, the 3150 115 118 118 sinusoidal vibration input level can be reduced 4000 113.5 116.5 116.5 5000 111 114.5 114.5 (notched) at the fundamental resonances of the 6300 107 110 110 hardmounted spacecraft or PAF/spacecraft assem- 8000 103 106 106 10000 100 103 103 bly to more realistically simulate flight loading OASPL 140.8 143.8 143.8 Duration 60 sec 120 sec 60 sec conditions. This has been accomplished on many Note: Delta II 7320 and 7325 acoustic environment is previous spacecraft in the lateral axis by correlating 2.0 dB lower M046,T4-10 one or several accelerometers mounted on the

Table 4-12. Sinusoidal Vibration Acceptance Test Levels Frequency Acceptance Sweep Axis (Hz) test levels rate Thrust 5 to 6.2 1.27 cm (0.5 inch) double amplitude 4 octaves/minute 6.2 to 100 1.0 g (zero to peak) Lateral 5 to 100 0.7 g (zero to peak) 4 octaves/minute M046,T4-13

Table 4-13. Sinusoidal Vibration Qualification Test Levels Frequency Acceptance Sweep Axis (Hz) test levels rate Thrust 5 to 7.4 1.27 cm (0.5 inch) double amplitude 2 octaves/minute 7.4 to 100 1.4 g (zero to peak) Lateral 5 to 6.2 1.27 cm (0.5 inch) double amplitude 2 octaves/minute 6.2 to 100 1.0 g (zero to peak) M046,T4-14

Table 4-14. Sinusoidal Vibration Protoflight Test Levels Frequency Acceptance Sweep Axis (Hz) test levels rate Thrust 5 to 7.4 1.27 cm (0.5 inch) double amplitude 4 octaves/minute 7.4 to 100 1.4 g (zero to peak) Lateral 5 to 6.2 1.27 cm (0.5 inch) double amplitude 4 octaves/minute 6.2 to 100 1.0 g (zero to peak) M046,T4-15 4-20 Md spacecraft to the bending moment at the PAF/ 4.2.5 Dynamic Analysis Criteria and Balance Requirements spacecraft separation plane. The bending moment is then limited by: (1)introducing a narrow-band 4.2.5.1 Two-Stage Missions. Two-stage mis- notch into the sinusoidal vibration input program, sions utilize the capability of the second stage to or (2)controlling the input by a servo-system using provide the terminal velocity, roll, final spacecraft a selected accelerometer on the spacecraft as the orientation, and separation. limiting monitor. A redundant accelerometer is Spin Balance Requirements. For nonspinning usually used as a backup monitor to prevent shaker spacecraft, there is no dynamic balance constraint, runaway. but the static imbalance directly influences the The Delta II program normally conducts a space- spacecraft angular rates at separation. When there craft/launch vehicle coupled dynamic loads analy- is a separation tipoff constraint, the spacecraft CG sis for various spacecraft configurations to define offset must be coordinated with the Delta Program the maximum expected bending moment in flight Office for evaluation. at the spacecraft separation plane. In the absence of Two-Step Separation System. For missions in a specific dynamic analysis, the bending moment is which there is a critical constraint on separation limited to protect the payload attach fitting, which tipoff angular rate, a two-step separation system is designed for a wide range of spacecraft configu- can be employed. The second stage and spacecraft rations and weights. The spacecraft user should are held together by loose fitting latches following consult the Delta Program Office before develop- primary separation of the nuts and bolts or clamp ing the sinusoidal vibration test plan for informa- bands. After a sufficient time (30 sec) for the tion on the spacecraft/launch vehicle coupled dy- angular rates to dissipate, the latches are released namic loads analysis for that special mission or and the second-stage retro thrust provides the re- similar missions. In many cases, the notched sinu- quired relative separation velocity from the space- soidal vibration test levels are established from craft. previous similar analyses. Second-Stage Roll Rate Capability. For some 4.2.4.4 Shock Testing. High-frequency pyro- two-stage missions, the spacecraft may require a technic shock levels are very difficult to simulate low roll rate at separation. The Delta second stage mechanically on a shaker at the spacecraft system can command roll rates up to 4 rpm (0.42 rad/s) level. The most direct method for this testing is to using control jets. Higher roll rates are also pos- use a Delta II flight configuration PAF/spacecraft sible; however, the accuracy is degraded as the rate separation system and PAF structure with func- increases. Roll rates higher than 4 rpm (0.42 rad/s) tional ordnance devices. Spacecraft qualification must be assessed relative to specific spacecraft and protoflight shock testing is performed by in- requirements. Significantly higher roll rates may stalling the in-flight configuration of the PAF/ require the use of a spin-table assembly. spacecraft separation system and activating the 4.2.5.2 Three-Stage Missions. Three-stage mis- system twice. Spacecraft shock acceptance testing sions employ a spin-stabilized upper stage. The is performed in a similar manner by activating the spin table, third-stage motor, PAF, and spacecraft PAF/spacecraft separation system once. combination are accelerated to the initial spin rate 4-21 Md prior to third-stage ignition by the activation of four mated as required for the mission-specific space- to eight spin rockets mounted on the spin table. Two craft characteristics. rocket sizes are available to achieve the desired Spin Rate Capability. Spin-up of the third stage/ spin rate. spacecraft combination is accomplished by activat- Spin Balance Requirements. To minimize the ing small rocket motors mounted on the spin table cone angle and momentum vector pointing error of that supports the payload. Spin direction is clock- the spacecraft/third-stage combination after wise, looking forward. Spin rates from 30 to 110 second-stage separation, it is necessary that the rpm are attainable for a large range of spacecraft roll imbalance of the spacecraft alone be within speci- moments of inertia (MOI) as shown in Figure 4-16. fied values. The spacecraft should be balanced to Nominal spin rates can be provided within ±5 rpm produce a CG within 1.3 mm (0.05 in.) of the for any value specified in the region of spin rate centerline, and a principal axis misalignment of less capability (Figure 4-16). Once a nominal spin rate than 0.25 deg with respect to the centerline. The has been determined, 3σ variations in relevant spacecraft centerline is defined as a line perpen- parameters will cause a spin rate prediction uncer- dicular to the separation plane of the spacecraft that tainty of ±15% about that nominal value at space- passes through the center of the theoretical spacecraft separation. craft/PAF diameter (refer to Section 5). Because orbit errors are dependent upon spin A composite balance of the entire third-stage/ rate, the magnitude of the orbit errors must be spacecraft assembly is not required. It has been assessed relative to the mission requirements and shown analytically that the improvements derived spacecraft mass properties before final resolution from a composite balance were generally small and of the spin rate for a specific spacecraft mission is do not justify the handling risk associated with accomplished. spacecraft spin balance on a live motor. Spin-Up Angular Acceleration. Spin-up of the For most spinning spacecraft, it has been demon- third-stage assembly imposes angular acceleration strated that the static and dynamic balance limits loads upon the spacecraft. The maximum angular defined herein can be satisfied. For the larger, acceleration that will occur while attaining a heavier spacecraft where such a constraint may be desired spin rate is fixed by spin motor thrust difficult to satisfy, the effects of broadened toler- characteristics. ances are analyzed on a per-case basis. The Delta II spin system utilizes two different The angular momentum/velocity pointing errors spin motors in various combinations to attain speci- and cone angle are highly dependent upon the fied spin rates. Figure 4-17 shows the maximum spacecraft spin rate, CG location, moments and angular acceleration that could be incurred by the products of inertia, NCS operation during system. Two curves are shown, one for a nominal upper-stage motor burn and coast periods, and the propellant temperature condition of 70°F (21.1°C) spacecraft energy dissipation sources during the and another for a maximum allowable temperature coast periods. The Delta Program Office, therefore, of 130°F (54.4°C) and +3σ burn rate. should be consulted if the above constraints cannot Figure 4-17 is based on the maximum motor be met. Pointing errors and cone angles are esti- thrust which occurs for a duration of approximately 4-22 Md H54875.1 120 AAAAAAAAAAAAAAAA 100AAAAAAAAAAAAAAAA AAAAAAAAAAAAAAAA 80AAAAAAAAAAAAAAAA

AAAAAAAAAAAAAAAARegion of

60AAAAAAAAAAAAAAAASpin Rate Capability Spin Rate (rpm) Rate Spin

40AAAAAAAAAAAAAAAA AAAAAAAAAAAAAAAA

0 100 200 300 400 500 600 700 Spacecraft Roll MOI (slug-ft2)

723 1446 2170 2893 3616 4340 5663 Spacecraft Roll MOI (kg-m-sec2) Figure 4-16. Delta II Spin Rate Capability

H54874

14AAAAAAAAAAAAAAAA AAAAAAAAAAAAAAAA

) 12 2 AAAAAAAAAAAAAAAA

10AAAAAAAAAAAAAAAA AAAAAAAAAAAAAAAA AAAAAAAAAAAAAAAA8 AAAAAAAAAAAAAAAA Angular Acceleration (rad/sec Acceleration Angular 6 AAAAAAAAAAAAAAAANormal AAAAAAAAAAAAAAAA4 AA+3-Sigma

2 30 40 50 60 70 8090 100 110

Spin Rate (rpm)

Figure 4-17. Maximum Expected Angular Acceleration Versus Spin Rate 4-23 Md 30 msec during ignition. If the maximum accelera- Table 4-15. Third Stage Mass Properties Before motor After motor tion is excessive, a detailed angular acceleration ignition burnout history can be provided for customer evaluation. If Weight (kg/lb) 2213/4878 191/422 CG aft of spacecraft not tolerable, special provisions such as sequential separation plane 780/30.7 808/31.8 firing of spin motors can be considered. (mm/in.) Spin MOI 385/284 45/33 Spacecraft Energy Dissipation During Coast (kgÊm2/slug-ft2) Transverse MOI Periods. Dissipation of spacecraft energy caused 454/335 92/68 (kgÊm2/slug-ft2) by nutation dampers, fuel slosh in the propellant M029, T036.2, 9/22/95, 9:21 AM tanks, flexible antennas, etc., can cause divergence in the cone angle between the spin axis of the upper stage and spacecraft and operates during the spacecraft/third-stage combination and its angular motor burn and postburn coast phase. momentum vector when the spin moment of inertia The NCS design concept uses a single-axis rate is less than the transverse moment of inertia. During gyro assembly (RGA) to sense coning and a mono- the periods between (1) third-stage separation and propellant (hydrazine) propulsion module to pro- ignition and (2) burnout and spacecraft separation, vide control thrust. The RGA angular rate signal is cone angle buildup can affect orbit accuracy, clear- processed by circuitry that generates thruster on/off ance between the spacecraft and the PAF during commands. Minor changes to the control circuitry separation, and spacecraft coning/momentum point- are required to accommodate different spacecraft ing after separation. mass properties and spin rates. The effect of energy dissipation is highly dependent The NCS nominal characteristics are listed in upon the mass properties and spin rate of the space- Table 4-16. Spacecraft weights less than 1500 lbs craft/third stage combination. In order for the Delta may require additional NCS modifications for the Program Office to evaluate the effect on a particular high third stage burnout acceleration. mission, the spacecraft agency must provide a worst -case energy dissipation time constant for the combined third stage and spacecraft for conditions Table 4-16. NCS Nominal Characteristics Propellant weight 2.72 kg/6.00 lb before and after third stage burn. Mass properties Helium prepressure 2.26 × 106 N/m2/400 psia for the third stage are shown in Table 4-15. Thrust 164.6 N/37 lb Minimum Isp (pulsing mode) 202.5 sec Nutation Control System. The NCS is designed Pressure at end of blowdown 9.7 x 105 N/m2/141 psia Transverse rate threshold 2 deg/sec to maintain small cone angles of the combined M029-T017.1

4-24 Md H54794 Section 5 SPACECRAFT INTERFACES Spacecraft This section presents the detailed descriptions and requirements of the mechanical and electrical interfaces of the launch vehicle with the spacecraft for two- and three-stage missions. PAF

5.1 DELTA THIRD-STAGE DESCRIPTION The third stage of the Delta II 7325 and 7925 Third Stage vehicles (Figure 5-1) consists of a Thiokol Star 48B Star 48B Motor solid-propellant rocket-motor, a cylindrical payload attach fitting (PAF) with a clamp assembly and Spin Table four separation spring actuators, a nutation control system (NCS), a spin table with bearing assembly, spin rockets, a clamp assembly and four separation spring actuators, an ordnance sequencing system, a telemetry system, and a yo-weight system for tum- bling the stage after separation. If required, a yo-yo DAC94092 despin system can be incorporated into the stack as Figure 5-1. Delta II Upper Stage a nonstandarad option in place of the yo-weight system to despin the spacecraft. The pre- and off-loaded condition. The amount of off-load is a post-burn mass properties of the stage are summa- function of spacecraft weight and the velocity re- rized in Table 4-15. quirements of the mission. The Star 48B motor has a diameter of 1244.6 mm Because of the development time and cost asso- (49.0 in.) and an overall length of 2032.0 mm (80.0 ciated with a custom PAF, it is to the advantage of in.) including an extended nozzle. The motor has the spacecraft agency to use the existing PAF de- two integral flanges, the lower for attachment to the signs. As early as possible in the design phase, third-stage spin table and the upper for attachment selection of an appropriate PAF should be coordi- to the 3712 PAF. The motor consists of a nated with the Delta Program Office. carbon-phenolic exit cone, 6AL-4V titanium 5.2 PAYLOAD ATTACH FITTINGS FOR high-strength motor case, silica-filled rubber insu- THREE-STAGE MISSIONS lation system, and a propellant system using In general, the component, sequencing, and sepa- high-energy TP-H-3340 ammonium perchlorate and ration system designs are the same for all three-stage aluminum with an HTPB binder. applications. The spacecraft is fastened to the PAF The Star 48B motor is available in propellant by a two-piece V-block-type clamp assembly, which off-loaded configurations. The motor is currently is secured by two instrumented studs, used for qualified for propellant weights ranging from 2010 clampband testing. Spacecraft separation is initi- kg (4430 lb) to 1739 kg (3833 lb) in the maximum ated by actuation of ordnance cutters that sever the 5-1 Md two studs. Clamp assembly design is such that Table 5-1. Notes Used in Configuration Drawings cutting either stud will permit spacecraft separa- 1. Interpret dimensional tolerance symbols in accordance with American National Standard Institute tion. Springs assist in retracting the clamp assembly (ANSI) Y14.5M-1982. The symbols used in this section are as follows: into retainers after release. A relative separation Flatness G velocity ranging from 0.6 to 2.4 m/s (2 to 8 ft/sec) Circularity E is imparted to the spacecraft by four spring actua- Parallelism // tors. Specific mission-oriented pads may be pro- Perpendicularity ⊥ (squareness) vided on the PAF at the separation plane to interface Angularity ∠ with spacecraft separation switches. A yo-weight Circular runout tumble system despins and imparts a coning motion Total runout to the expended third-stage motor 2 sec after True position ⊕ spacecraft separation to change the direction of its Concentricity momentum vector and prevent recontact with the Profile of a surface spacecraft. Diameter ∅ 2. Unless otherwise specified, tolerances are as PAF components are mounted on its surface. All follows: hardware necessary for mating and separation (e.g., Decimal mm 0.X = ±0.76 PAF, clamp assembly, studs, explosives, and tim- 0.XX = ±.0.03 in. 0.XX = ±0.03 ers) remains with the PAF upon spacecraft separa- 0.XXX = ±0.015 tion. Table 5-1 applies to the various configuration Angles = ±0 deg. 30 min 3. Dimensions apply at 69°F (20°C) with interface in drawings that accompany this section. unrestrained condition. 125 The 3712 PAF shown in Figure 5-2 is the 4. All machine surface roughness is per ANSI B46.1, 1985. interface between the upper-stage motor and the 5. The V-block/PAF mating surface is chemically conversion-coated per MIL-C-5541, Class 3 spacecraft. The PAF is approximately 305 mm (12 M029-T012-4/4/96-2:06 PM in.) high and 940 mm (37 in.) in diameter. It supports the clamp assembly that attaches the space- configurations and their associated spacecraft craft to the upper stage and allows the spacecraft to interface requirements are shown in be released at separation. It provides mounts for the Figures 5-4 through 5-16. four separation spring actuators, two electrical dis- The estimated capabilities shown in Figure 5-3 connects (if applicable), event sequencing system, are based on extrapolation of analytical results for upper-stage telemetry, and an NCS. numerous generic spacecraft designed to the spacecraft stiffness requirements specified in Section The 3712 PAF is available with three forward 4.2.3.2. The capability for a specific spacecraft flange configurations, designated 3712A, B, and C. (with its own unique mass, size, flexibility, etc.) The maximum clamp assembly preload for these may vary from that presented in Figure 5-3. There- configurations is given in Table 5-2. Figure 5-3 provides estimated capabilities of Table 5-2. Maximum Clamp Assembly Preload these configurations on the 7925 launch vehicle for Spacecraft PAF Max flight flange angle each of two Delta fairings. The capabilities are PAF preload (N/lb) (deg) plotted in terms of spacecraft weight and CG location 3712A 30,248/6800 15 3712B 25,355/5700 20 above the separation plane. The flange 3712C 25,355/5700 20 T037 5-2 Md DAC123579 H54797

Figure 5-2. 3712 PAF Assembly 02738REU7

100 2.5 Note: The capability is provided as a Note: guide for spacecraft design and is 90 Note: subject to verification by coupled Note: loads analysis. The levels shown Note: are for a 10-ft-diameter metal fairing With 9.5-ft-dia Fairing Note: and are conservative for the 10-ft-diameter 80 Note: composite fairing. These levels will be 2.0 Note: updated in a future revision.

60 1.5

CG Distance From Separation Plane (in.) 40 With 10-ft-dia Fairing 1.0 CG Distance From Separation Plane (m)

30 Allowable Region

20 0.5 2.02.5 3.03.5 4.0 4.5 5.0 103 lb

1.0 1.5 2.0 103 kg Spacecraft Weight

Figure 5-3. Capability of the 3712 PAF Configuration on 7925 Vehicle (Download Figure) 5-3 Md 02881REU7

III

22° 30' 12° 30' Bolt Cutter (2 Places)

Battery

Ordnance Sequencing System Panel

Coning Control Clamp Band Retainer (10 Assembly Places)

NCS Thruster Arm

¯ 32.50

Keyway IV II

4 x 45° 0' 2.5 deg Propellant/ Pressurant Clamp Assembly Tank

¯ 48.00 Telemetry Control Box

Spring Actuator (4 Places)

S/C Electrical Disconnect Rate Gyro Bracket (2 Places) I

View Looking Aft

Detail A (See Figures 5-5, 5-8, and 5-11)

304.80 12.00

939.80 ¯ 37.00

Side View of 3712 PAF Without Mounted Components

Figure 5-4. 3712 PAF Detailed Assembly (Download Figure) 5-4 Md 945.26 ± 0.076 02746REU7 +0.00 ¯ 37.215 ± 0.003 939.8 Ð0.13 Do Not Break ¯ 0.25 16.0 37.000+0.000 Sharp Edges 0.13 0.63 0.051/0.002 A Ð0.005 2 x R 2.36 0.010 1.52 2.21 -A- 0.005 1.40 ± 0.093 ° 3.3 0.64 0.12 15 0' 0.060 R 0.13 0.025 ± 0.005 0.087 ± 0° 15' 0.055 6.35 0.250 0.025/0.001

16.0 5.84 ± 0.076 0.63 0.230 ± 0.003 45 deg R

6.35 3.56 0.250 Chemical Conversion 0.140 63 7.6 Coat Per MIL-C-5541, 7.6 3 x R Class 3 0.30 0.050 9.53 0.375 mm in. 20.57 0.81 Detail A From Figure 5-4 Figure 5-5. 3712A PAF Detailed Dimensions (Download Figure)

D See Figure 5-7 02747REU7 Spacecraft

Separation -B- Plane 876.30 ± 0.254 0.050/0.002 A

¯ 34.500 ± 0.010 PAF 945.26 ± 0.076 ¯ 37.215 ± 0.003

Section A-A 0.050/0.002 A C 4 x ¯ 50.8/2.00 Area IV See Figure 5-7 MS 3464E37-50S for 340-lbf Separation Spring Electrical connector on Spacecraft Side (Typ 2 Places) ¯ 0.245/0.010 M B C S (Area extends from the separation plane and forward 7.11/0.280) See Figure 5-7 B B 1219 ¯ 48.000 Keyway on 825.50 Outboard side ¯ 32.500 (Typ 2 Places) III I 22° 30' A A 4 x 45 deg

2 x Ø 182.88/7.20 Area for Spacecraft/PAF Electrical Connectors (Area extends from the separation plane and forward 50.8/2.00) mm ¯ 0.762/0.030 S BCS in. II View Looking Forward

Figure 5-6. 3712A PAF Spacecraft Interface Dimensional Constraints (Download Figure) 5-5 Md 02750REU7 mm 30 deg ± 0°30' in. +0.13 5.54 For Section Marked ° -0.00 60 0' 2 2 ± +0.005 Area = 492 mm /0.763 in. 15% + 0° 15' 0.218 -0.000 4 4 ± - 0° 0' I = 45,784 mm /0.110 in. 15% Chord Line Applicable Length, L = 25.4 mm/1.0 in.

0.25 ± 0.13 2 x R +0.25 0.010 ± 0.005 17.48 -0.00 939.80 ¯ +0.010 37.000 0.688 -0.000 View C From Figure 5-6 9.7

0.38 7.6 0.30 1.27 R 0.050 5.84 ± 0.076 30 deg 0.23 ± 0.003 3.56 0.140 L

15 deg ± 0° 15'

Ð B Ð

R 1.27/0.050 0.03/0.001 E Over 16.0/0.63 16 Wide Surface 0.63 945.26 ± 0.076 ¯ 37.215 ± 0.003 63 ÐAÐ Chemical Conversion Coat per MIL-C-5541, Class 3 +0.13 940.94 -0.00 View D ¯ +0.005 From Figure 5-6 37.045 - 0.000

+0.13 3.56 -0.00 +0.005 0.33/0.013 0.140 -0.000 ÐCÐ +0.13 0.254 -0.00 0.76 +0.13 2 x R 1.78 +0.005 0.03 -0.00 0.010 -0.000 +0.005 R 0.38/0.015 Full Relief 0.070 -0.000 0.64/0.025 Deep

7.11 23.26 0.280 ¯ +0.13 0.916 0.13 -0.00 ÐBÐ 2 x R 25.81 3.81 ¯ +0.005 1.016 0.150 0.005- 0.000 48.67 ¯ min View E 1.916 Section B-B From Figure 5-6

Figure 5-7. 3712A PAF Spacecraft Interface Dimensional Constraints (Views C, D, E, and Section B-B) (Download Figure) 5-6 Md 958.85 ± 0.076 02744REU7 Do Not Break Sharp Edges ¯ 0.050/0.002 A mm 37.750 ± 0.003 in. 11.43 0.45* 954.02 + 0.000 Ð 0.127 ¯ Ð A Ð 3.23 ± 0.076 37.560 + 0.000 Ð 0.005 2.36 0.127 ± 0.003 2.21 0.13 876 ¯ 0.381/0.015 A 0.093 0.25 34.50 0.087 2 X R 20° 0' 0.005 5.33 ± 0° 15' 0.010 0.210 45° ± 5° × 0.12

0.025/0.001 1.52 16 Applicable over 0.45 dim 1.40 0.63 as noted * 0.060 63 0.055 7.87 0.31 Chemical 7.62 9.7 Conversion R 0.300 0.38 Coat per MIL-C-5541, 4.60 Class 3 2.29 0.181 R 0.090 3.1 940.05 R ¯ ¯ 0.254/0.010 s A s 0.12 37.010 Detail A From Figure 5-4 Figure 5-8. 3712B PAF Detailed Dimensions (Download Figure) D See Figure 5-10 02749REU7 mm

in. Spacecraft 0.050/0.002 A

Separation -B- Plane 876.30 ± 0.254 ¯ PAF 34.500 ± 0.010 958.85 ± 0.076 ¯ 37.750 ± 0.003 Section A-A 0.050/0.002 A C IV See Figure 5-10 ¯ 0.254/0.010 M B C S MS 3464E37-50S Electrical Connector 4 x ¯ 50.8/2.00 Area for on Spacecraft Side 340-lbf Separation Spring (Typ 2 Places) (Area Extends from the Separation Plane and Forward 7.11/0.280)

See Figure 5-10 B B 1219 Keyway on ¯ 48.000 Outboard Side (Typ 2 Places)

III I 22° 30' 825.50 A ¯ 32.500 A

4 x 45 deg 2 x ¯ 182.88/7.20 Area for Spacecraft/PAF Electrical Connectors (Area Extends from the Separation Plane and Forward 50.8/2.00)

¯ 0.762/0.030 S BCS II View Looking Forward Figure 5-9. 3712B PAF Spacecraft Interface Dimensional Constraints (Download Figure) 5-7 Md 02753REU7

mm 30 ° ±0° 30' in. +0.13 5.54 -0.00 For Section Marked 60° 0' 2 2±15% ° +0.005 Area = 269 mm /0.417 in. +0 15' 0.218 4 -0 ° 0' -0.000 I = 11,654 mm /0.028 in. 4 ±15% Chord Line Applicable Length, L = 25.4 mm/1.0 in.

0.25 ± 0.13 2 x R ± +0.25 0.010 0.005 17.48 -0.00 0.688 +0.010 -0.000 939.80 View C ¯ 37.00 From Figure 5-9

3.05 R 6.35 0.12 R 0.25

± 3.22 0.070 20 ° 7.6 L 0.127 ± 0.03 ±0° 15' 0.30

0.70 R 0.03 0.03/0.001 - B -

E +0.51 5.08 -0.00 +0.020 2.28 ±0.25 0.200 R - 0.000 0.090 ±.010 +0.13 937.87 -0.00 ¯ +0.005 63 Chemical Conversion Coat 36.924 - 0.000 per MIL-C-5541, Class 3 View D 0.33/0.013 -A- From Figure 5-9 +0.13 955.17 -0.00 ¯ +0.005 37.605 - 0.000 +0.13 3.56 -0.00 0.76 R R 0.38/0.015 Full Relief -C- +0.005 0.03 0.140 -0.000 0.64/0.025 Deep +0.13 1.78 -0.00 +0.13 0.254 -0.00 0.070 +0.005 2 x -0.000 0.010 +0.005 -0.000 7.11 23.26 0.280 ¯ 0.916 - B - 25.81 ¯ 3.81 0.13 +0.13 1.016 0.150 2 x R -0.00 48.67 ¯ min 0.005 +0.005 - 0.000 1.916 View E Section B-B From Figure 5-9

Figure 5-10. 3712B PAF Spacecraft Interface Dimensional Constraints (Views C, D, and E and Section B-B) (Download Figure) 5-8 Md 02755REU7 mm 2.36 954.02 +0.000/Ð0.127 2.21 ¯ ÐAÐ in. 37.560 +0.00/Ð0.005 0.093 937.108 +0.000/Ð0.254 0.087 ¯ 0.127/0.005 A 958.85 ± 0.076 36.894 +0.000/Ð0.010 ¯ 0.050/0.002 A 0.25 37.750 ± 0.003 0.13 2 X R 1.52 0.010 20° 0' 0.005 63 Chemical Conversion ±0° 15' 1.27 Coat Per MIL-C-5541 0.060 +0° 0' Class 3 Gold Do Not Break Sharp 44 ° 0.050 Ð0° 30' Edges 6.35 0.250 0.025/0/0.001 ÐBÐ 3.23 ± 0.076 R 9.7 4.06 0.127 ± 0.003 0.38 0.160 2.29 7.87 R 0.090 0.31 ¯ 901.45 0.381/0.015 A Detail A R 3.1 35.490 From Figure 5-4 0.12 ¯940.05 ¯ 0.254/0.010 S A S 37.010 Figure 5-11. 3712C PAF Detailed Dimensions (Download Figure) 02756REU7 mm D See Figure 5-13 in. Spacecraft

Separation -B- Plane 0.050/0.002 A 901.70 ± 0.254 ¯ 35.500 ± 0.010 PAF 958.85 ± 0.076 ¯ 37.750 ± 0.003 Section A-A 0.050/0.002 A

C See Figure 5-13 IV

¯ 0.254/0.010 B C S MS 3464E37-50S Electrical Connector 4 x ¯ 76.2/3.00 Area for on Spacecraft Side 200-lbf Separation Spring (Area Extends from the (Typ 2 Places) Separation Plane and Forward 7.11/0.280)

See Figure 5-13 B B ¯ 1219 Keyway on 48.000 Outboard Side (Typ 2 Places)

III I ° 22 30' ¯ 825.50 A 32.500 A

4 x 45 deg

2 x ¯ 182.88/7.20 Area for Spacecraft/PAF Electrical Connectors (Area Extends from the Separation Plane and Forward 50.8/2.00)

¯0.762/0.030 S BCS II View Looking Forward Figure 5-12. 3712C PAF Spacecraft Interface Dimensional Constraints (Download Figure) 5-9 Md 02757REU7 mm 30 °± 0° 30' in. +0.13 5.54 -0.00 For Section Marked ° 60 0' +0.005 Area = 269 mm 2 /0.417 in. 2 ± 15% + 0 ° 15' 0.218 -0.000 4 4± - 0° 0' I = 11,654 mm /0.028 in. 15% Chord Line Applicable Length, L = 25.4 mm/1.0 in.

0.25 ± 0.13 +0.25 939.80 2 x R ± 17.48 ¯ 0.010 0.005 -0.00 37.00 +0.010 0.688 View C -0.000 From Figure 5-12 3.05 R 0.12 6.35 R 7.6 0.25 2.28 ± 0.25 R 0.30 L 0.090 ±.010

-B- 20°± 0 °15' 0.03 0.001

E +0.51 ± 0.70 5.08 3.22 0.070 R -0.00 0.127 ± 0.003 0.03 +0.020 47 °± 0°30' 0.200 - 0.000

937.87 ± 0.254 ¯ ± 63 36.924 0.010 Chemical Conversion Coat per MIL-C-5541, Class 3

View D From Figure 5-12 ÐAÐ

955.17 +0.13 -0.00 ¯ 37.605 +0.005 - 0.000 R 0.38/0.015 Full Relief +0.13 0.64/0.025 Deep 3.56 0.76 -0.00 R 0.03 3.81 +0.005 0.33/0.013 0.140 0.150 -0.000 -C- 0.254 +0.13 2 x -0.00 1.78 +0.13 -0.00 0.010+0.005 7.11 40.89 +0.005 -0.000 0.280 ¯ 0.070 -0.000 1.610 - B - 45.97 1.810

74.7 ¯ 2.94 0.13 +0.13 2 x R -0.00 Section B-B +0.005 View E 0.005 From Figure 5-12 - 0.000

Figure 5-13. 3712C PAF Spacecraft Interface Dimensional Constraints (Views B-B, C, D and E) (Download Figure) 5-10 Md 02758REU7

III ° 22° 30' Bolt Cutter 12 30' (2 Places) See Figure 5-15 See Figure 5-15 B Battery C

Ordnance Sequencing System Panel

C B Clamp Band Retainer (10 Coning Control Places) Assembly

NCS Thruster Arm

¯ 32.50 Keyway IV II

4 x 45° 0' 2.5 deg

Clamp Assembly NCS Tank

¯ 48.00

Telemetry Control Box Spring Actuator (4 Places) S/C Electrical Disconnect Bracket D Rate Gyro (2 Places) See Figure 5-16 I View AÐA

Spacecraft

PAF 3rd Stage Rocket Motor

Figure 5-14. 3712 PAF Interface (Download Figure) 5-11 Md 02759REU7 Clamp Assembly

Spacecraft mm Clamp Retainer in.

PAF (3712A Shown) View C-C From Figure 5-14 (Rotated 25-deg CW) 2.79 (Max) 0.110

1219 ¯ 48.00

Spacecraft Connector Mounting Panel +1.4/0.055 ( 18.29/0.720 ) Clamp Retainer -0.38/0.015 Clamp Assembly

Payload Ring Spring Pad 46.88/1.846 +0.0/0.0 -0.05/0.002 Separation Separation Plane -B- Plane

65.18 ± .38 2.566 0.015

Separation Springs

Spacecraft Electrical Disconnect Bracket

Balance Weights

3712 PAF

Rocket Motor

View B-B From Figure 5-14 (Rotated 45-deg CW)

Figure 5-15. 3712 Clamp Assembly and Spring Actuator (Download Figure) 5-12 Md 02760REU7

Clamp Assembly Bolt Cutter Bracket

Bolt Cutter Shield V-Block

End Fitting Calibrated Stud

Bolt Cutter Bracket Bolt Cutter Shield

Calibrated Stud E Explosive Bolt Cutter View E-E

View D From Figure 5-14 (View Rotated 90-deg CW)

Figure 5-16. 3712 PAF Bolt Cutter Detailed Assembly (Download Figure) fore, when the spacecraft configuration is deter- ered. Separation spring forces ranging from 890 to mined, Delta Program Office will initiate a coupled 1512 N (200 to 340 lb) per spring are available. loads analysis to verify that the structural capability The Delta program has experience utilizing non- of the launch vehicle is not exceeded. standard spring forces up to 1700 N (400 lb); spring For a spacecraft that requires a longer PAF in forces outside the standard offering may be negoti- order to eliminate its interference with the upper ated with the Delta Program Office. stage, an optional cylindrical extension adapter 5.3 PAYLOAD ATTACH FITTINGS FOR with customized length can be inserted between the TWO-STAGE MISSIONS PAF and the upper stage. The extension adapter Delta offers several PAF configurations for use reduces the spacecraft allowable CG capability on two-stage missions. Selection of an appropriate presented in Figure 5-3 approximately by the length PAF should be coordinated with the Delta Program of the adapter. Office as early as possible in the spacecraft design Although the above discussion provides a guide, phase. the actual selection of the PAF configuration is The PAF for two-stage missions has a separation made after discussions between the Delta Program system that is activated by a power signal from the Office and the spacecraft contractor. During these second stage, rather than by a self-contained com- discussions, separation requirements are also consid- ponent, as on the three-stage PAFs. 5-13 Md On two-stage Delta II configurations, the space- machined from an aluminum forging. Since this craft is separated by the activation of explosive nuts fitting was designed specifically to interface with (for the 6019 and 6915 PAFs) or by the release of a the NASA Multimission Modular Spacecraft V-band clamp assembly (for the 5624 and 6306 (MMS), users should coordinate with the Delta PAFs) followed by the action of the second-stage, Program Office to ensure that the required interface helium-gas retro system. An optional secondary stiffness is provided. latch system is available that will retain the space- Figure 5-18 provides the estimated capability of craft until the energy introduced by the separation the PAF on a 7920 launch vehicle in terms of process is damped out prior to the action of the retro spacecraft weight and CG location above the sepa- system. Some PAFs come standard with latch ration plane for two payload fairing configurations. systems, while the 5624 is not available with latches. The estimated capability is based on extrapolation (See Sections for each PAF.) The relative separa- of analytical results for generic spacecraft de- tion velocity is approximately 0.3 m/s (1 ft/sec). signed to the spacecraft stiffness requirements specified in Section 4.2.3.2. The capability for a 5.3.1 The 6019 PAF Assembly specific spacecraft (with its own unique mass, size, The 6019 PAF assembly, shown in Figure 5-17, flexibility, etc.) may vary from that presented in is approximately 483 mm (19 in.) high and 1524 Figure 5-18. Therefore, after the spacecraft con- mm (60 in.) in diameter. The one-piece structure is figuration is determined, the Delta Program Office

H54796

Figure 5-17. 6019 PAF Assembly 5-14 Md 02739REU7

120 3 Note: The capability is provided as a 110 guide for spacecraft design and is subject to verification by coupled loads analysis. The levels shown are 100 for a 10-ft-diameter metal fairing and are conservative for the 10-ft-diameter composite fairing. These levels will be updated in a future revision. 90

80 2 With 9.5-ft-dia Fairing

60 With 10-ft-dia Fairing CG Distance From Separation Plane (in.) CG Distance From Separation Plane (m) Allowable Region 50

40 1 567891011103 lb

3 4 5103 kg Spacecraft Weight

Figure 5-18. Capability of the 6019 PAF on 7920 Vehicle (Download Figure) will initiate a coupled loads analysis to verify that spacecraft. For the 6019 PAF a two-step release the structural capability of the launch vehicle is not system consisting of the three explosive nuts and a exceeded. secondary latch system is employed. This requires The base of the fitting is attached to the forward that a small lip (MDC-provided) be installed on the ring of the second stage. The spacecraft is fastened spacecraft at each of the attach bolt locations to the 1524-mm (60 in.) PAF mating diameter at (Figures 5-19, 5-21, and 5-22). Following activation of the explosive nuts, three retaining latches are three equally spaced hard points with 15.9-mm released and the second stage is retro-fired so that (0.625 in.) bolts that are preloaded to 53,380 N there is a minimal tip-off of the spacecraft. (12,000 lb) tension. Separation of the spacecraft from the launch vehicle occurs when three explo- 5.3.2 The 6915 PAF Assembly sive nuts are activated and the second stage is The 6915 PAF assembly (Figure 5-23) is ap- backed away by the helium retro system. proximately 381 mm (15 in.) high and 1743 mm The spacecraft interface is shown in Figures (68.6 in.) in diameter. This one-piece structure is 5-19 and 5-20. It should be noted that the launch machined from an aluminum forging. vehicle requires access on the spacecraft side of the Figure 5-24 provides the estimated capability of separation plane for installation of the three attach the PAF on the 7920 launch vehicle in terms of bolts and bolt catcher assemblies. Upon separation, spacecraft weight and CG location above the sepa- the bolts and catcher assemblies are retained by the ration plane for two payload fairing configurations. 5-15 Md 02761REU7

mm in. A 88.9 3.50 Spacecraft Dia

Section A-A A (Typ 3 Places) 203.2 8.00 Side View of 6019 PAF Bolt Catcher Envelope (Required for Installation) Bolt Catcher 1524.00 ¯ (Ref) 60.00 Secondary Latch System Spacecraft Bracket—MDA-Supplied

Section B-B B B Shown Below Separation 0.127/0.005 Plane 17.42/0.686 15.189/0.598 34.943/1.3757 ¯17.47/0.688 ¯ 15.240/0.600 34.950/1.3760 -B- ¯ 0.025/0.001 ¯ 0.127/0.005 -B- 32 ¯ 0.025/0.001 60° 0' (Ref)

M45932/1-9CL Insert, 2 Required Chemical (for 4.76/0.1875 Dia Bolt) Conversion Coat 10.52/0.414 Tap Drill Depth per MIL-C-5541, 125 ¯ 0.711/0.028 Class 3 (For Secondary Latch System) 2.54 0.10 24.13 M45932/1-21CL Insert, 2 Required 0.950 (for 9.53/0.375 Dia Bolt) 18.29/0.720 Tap Drill Depth 14.22 63.50 0.56 ¯ 1.575/0.062 2.50 28.57/1.125 (Min)

50.80 (Typ) 2.00

27° 54' ± 0° 15' 25.40 (Typ) 1.00 15.24 Note: Matched 0.600 63.50/2.500 tooling for drilling interface holes, 31.78/1.250 tolerance within 11.10 R 0.437 46.99 ¯ 0.127/0.005 (Max) 1.850 of tooling

Section BÐB

Note: Constraints are the responsibility of the spacecraft agency

Figure 5-19. 6019 PAF Spacecraft Interface Dimensional Constraints (Download Figure) 5-16 Md 02762REU7

mm in. Matched Tooling I Provided for S/C Interface Hole Pattern II

1 3

7° 59'

120° 0' 120° 0'

1524.00 ¯ 60.000

III

Separation Plane 0.254/0.010 ÐAÐ 0.127/0.005

487.68 19.20

ÐAÐ

Figure 5-20. 6019 PAF Detailed Assembly (Download Figure) 5-17 Md 02763REU7 02764REU7 mm in.

34.793 Bolt Catcher 34.666 ¯ (Prior to Dry Lube) 1.3648 MDA Provided 1.3698 Attach Hardware Attach Bolt (Typ 2 Places)

Lockwire Catcher 17.437/0.6865 Assy as Shown Grounding ¯ Provisions 17.488/0.6885 MDA-Provided Spacecraft (Ref) Spacecraft Latch Interface Separation Plane Latch Dry Lube Per MIL-L-8937

Optional 10.16/0.400 Secondary Separation Separation Latch System Nut Plane

° Latch 60°+0 15' Mechanism - 0° 00'

Separation Nut

Section A-A (Typ 3 Places for 6019) Section A-A

A A A

A Side View

Top View

Figure 5-21. 6019 PAF Spacecraft Assembly (Download Figure 5-22. 6019 PAF Detailed Dimensions (Download Figure) Figure) 5-18 Md DAC115974 H54795

Figure 5-23. 6915 PAF 02741REU7

120 3 Note: The capability is provided as a 110 guide for spacecraft design and is With 9.5-ft-dia Fairing subject to verification by coupled loads analysis. The levels shown are for a 10-ft-diameter metal fairing and are 100 conservative for the 10-ft-diameter composite fairing. These levels will be updated in a future revision 90

80 2

70 With 10-ft-dia Fairing

60 CG Distance from Separation Plane (in.) Allowable Region CG Distance from Separation Plane (m)

40 1 3 56789101110 lb

3 3 4 510 kg Spacecraft Weight

Figure 5-24. Capability of the 6915 PAF on 7920 Vehicle (Download Figure) 5-19 Md The estimated capability is based on extrapolation 5.3.3 The 6306 PAF Assembly of analytical results for generic spacecraft designed The 6306 PAF assembly (Figures 5-30 through to the spacecraft stiffness requirements specified in 5-34) is approximately 152.4 mm (6 in.) high and Section 4.2.3.2. The capability for a specific space- 1600 mm (63 in.) in diameter. This one-piece craft (with its own unique mass, size, flexibility, structure is machined from an aluminum forging. etc.) may vary from that presented in Figure 5-24. Figure 5-35 provides the estimated capability of Therefore, when the spacecraft configuration is the PAF on the 7920 launch vehicle in terms of determined, the Delta Program Office will initiate a spacecraft weight and CG location above the sepa- coupled loads analysis to verify that the structural ration plane for different payload fairing configura- capability of the launch vehicle is not exceeded. tions. The estimated capability is based on extrapo- The base of the PAF is attached to the forward lation of analytical results for generic spacecraft ring of the second stage. Matched tooling is pro- designed to the spacecraft stiffness requirements vided for the hardpoint interface and the mechani- specified in Section 4.2.3.2. The capability for a cal/elecctrical interfaces, if required. The space- specific spacecraft (with its own unique mass, size, craft is fastened to the 1742.6-mm (68.6-in.) PAF flexibility, etc.) may vary from that presented in mating diameter at four equally spaced hard points Figure 5-35. Therefore, when the spacecraft con- with 15.9-mm (0.625-in.) bolts that are preloaded to figuration is determined, the Delta Program Office 53,380-N (12,000-lb) tension. The separation of the will initiate a coupled loads analysis to verify that spacecraft from the launch vehicle occurs when the structural capability of the launch vehicle is not four explosive nuts are activated and the second exceeded. stage is backed away by spring actuators. The base of the PAF is attached to the forward The spacecraft interface is defined in Figures ring of the second stage. The spacecraft is fastened 5-25 through 5-29. Matched tooling is provided for the hardpoint interface and the meachanical/electri- to the 1600 mm (63 in.) PAF mating diameter with cal interfaces, if required. It should be noted that the a V-band clamp assembly that is preloaded to 34,250 launch vehicle requires access on the spacecraft N (7,700 lb). Separation of the spacecraft from the side of the separation plane for installation of the launch vehicle occurs when the V-band clamp as- four attach bolts and bolt catcher assemblies. Upon sembly is released and the second stage is backed separation, the bolts and catcher assemblies are away by the helium retro system. A two-step retained by the spacecraft. For spacecraft systems release system is employed, which requires the requiring a minimal tip-off rate at separation, the addition of four holes in the spacecraft interface spring actuators are removed and a two-step release ring (Figure 5-36) to mate with PAF-mounted lat- system consisting of four explosive nuts and a eral restraints. Following release of the V-band secondary latch is employed. Use of this system clamp, the interaction between the spacecraft and requires that a small lip (MDA-provided) be added the retaining latches, structure, and damping springs to the spacecraft at each of the attach bolt locations settle the spacecraft rates (damping is provided by (Figures 5-27 and 5-28). Following activation of spacecraft-provided separation springs or launch the four explosive nuts, the four retaining latches vehicle-provided damping devices). Thirty seconds are released and the second stage is retro-fired. This later, the three retaining latches are retracted and the system limits the tip-off imparted to the spacecraft. second-stage retro system is activated backing the 5-20 Md 02765REU7

mm in

IV I

1930.4 ¯76.000 B B See Figure 5-29

1742.19 Matched Tooling Provided for Spacecraft Interface Hole Pattern ¯ 68.590

22.238 22.212 4 X ¯ 0.8755 0.8745 ¯ 0.130/0.005 M

III

II 39° 45'

See Figure 5-29 Actuator Support A (4 Places) -C- 0.130/0.005 A (4 Surfaces) 381.0 15.000 Electrical Bracket (2 Places) -A- A 1630.8 ¯ 64.205 -B-

Figure 5-25. 6915 PAF Detailed Assembly (Download Figure) 5-21 Md 02766REU7 mm in. A 89.9 3.50 Dia Spacecraft A Section A-A (Typ 4 Places) 203.2 8.00 Side View of 6915 PAF Bolt Catcher Envelope (Required for Installation) Bolt Catcher

1742.19 ¯ (Ref) Spacecraft 68.607 Secondary Latch System Bracket—MDA-Supplied

Section B-B Shown Below

BB Separation 0.127/0.005 Plane 17.42/0.686 15.189/0.598 ¯ 17.47/0.688 15.240/0.600 34.943/1.3757 ¯ 34.950/1.3760 -B- ¯ 0.025/0.001 ¯ 0.127/0.005 B 32 60° 0' ¯ 0.025/0.001 (Ref)

M45932/1-9CL Insert, 2 Required Chemical (for 4.76/0.1875 Dia Bolt) Conversion Coat 10.52/0.414 Tap Drill Depth per MIL-C-5541, 125 Class 3 ¯ 0.711/0.028 2.54 24.130 (For Secondary Latch System) 0.10 0.950 M45932/1-21CL Insert, 2 Required (for 9.53/0.375 Dia Bolt) 18.29/0.720 Tap Drill Depth 14.22 63.50 0.56 ¯ 1.575/0.062 2.50 28.57/1.125 (Min)

50.80 (Typ) 2.00

27° 54' ± 0° 15' 25.40(Typ) 1.00 15.24 Note: Matched 0.600 63.50/2.500 tooling for drilling interface holes, 31.78/1.250 tolerance within 11.10 R ¯ 0.127/0.005 0.437 46.99 of tooling (Max) 1.850

Section BÐB

Note: Constraints are the responsibility of the spacecraft agency

Figure 5-26. 6915 PAF Spacecraft Interface Dimensional Constraints (Download Figure) 5-22 Md 02768REU7 02767REU7 34.793 mm Attach Bolt 34.666 in ¯ (Prior to 1.3648 Dry Lube) 1.3698 17.437/0.6865 Bolt Catcher ¯ 17.488/0.6885 MDA-Provided Spacecraft Lockwire Latch Interface Dry Lube Per Catcher MIL-L-8937 Assembly as Latch Shown Mounting Hardware (Typ 2 Places) Provided by MDA 10.16/0.400

Spacecraft (Ref) Separation Separation Plane Plane ° ° +0 15' 60 0' -0° 0' Grounding Provisions

Latch Mechanism Separation Nut

Section A-A (Typical 4 Places)

Optional Secondary Latch System

Section A-A

Figure 5-27. 6915 PAF Spacecraft Assembly (Download Figure 5-28. 6915 PAF Detailed Dimensions (Download Figure) Figure) 5-23 Md 02769REU7.1

Male Separation Cone

PAF Actuator Support

View A-A (From Figure 5-25)

R 0.38/0.015 Full Relief 0.64/0.025 Deep Spacecraft 3.81 0.150 8.89 0.350 Spring Seal

23.27 ¯ 0.916 Separation Actuator Assembly Plane 25.81 1.016 Umbilical 48.67 ¯ Minimum Bracket 1.916

Actuator Support

PAF

View B-B (4 Places) (From Figure 5-25)

Figure 5-29. Actuator Assembly Installation – 6915 PAF (Download Figure) 5-24 Md H54659

Figure 5-30. 6306 PAF Assembly second stage away from the spacecraft. ‘The second Figure 5-42. Therefore, when the spacecraft con- stage then performs a contamination and collision figuration is determined, the Delta Program Office avoidance maneuver (CCAM) to remove the stage will initiate a coupled loads analysis to verify that from the vicinity of the spacecraft, followed by a the structural capability of the launch vehicle is not propellant depletion burn to safe the stage. exceeded. 5.3.4 The 5624 PAF Assembly The base of the PAF is attached to the forward The 5624 PAF assembly (Figures 5-37 through ring of the second stage. The spacecraft is fastened 5-41) is approximately 609.6 mm (24 in.) high and to the 1422.4 mm (56 in.) PAF mating diameter 1422.4 mm (56 in.) in diameter. This one-piece with a V-band clamp assembly that is preloaded to structure is machined from an aluminum forging. 17,350 N (3900 lb). The design of this PAF does Figure 5-42 provides the estimated capability of not allow for a spacecraft side latch. Separation of the PAF on the 7920 launch vehicle in terms of the spacecraft from the launch vehicle occurs spacecraft weight and CG location above the sepa- when the V-band clamp assembly is released and ration plane for different payload fairing configura- four spring actuators impart a relative separation tions. The estimated capability is based on extrapo- velocity between the spacecraft and the launch lation of analytical results for generic spacecraft vehicle. designed to the spacecraft stiffness requirements specified in Section 4.2.3.2. The capability for a 5.4 NEW PAFs specific spacecraft (with its own unique mass, size, New PAFs are currently being studied. The flexibility, etc.) may vary from that presented in design of these PAFs takes into account the use of 5-25 Md IV 02770REU7.1

CL Keyway

48° 0'

112° 30' 45° 30'

See Below A A

I III

122° 0'

CL Secondary Latch System (3 Places If Required) (See Figure 5-36)

1604.721 ¯ 63.178 1599.057 ¯ 62.955 75.9 B See Figure 5-32 2.99

12.192 0.48

152.4 6.00

9.779 0.385

27.9 1.10 Section A-A 1524.457 ¯ 60.018

Figure 5-31. 6306 PAF Detailed Dimensions (Download Figure) 5-26 Md 02771REU7.1 mm in.

1604.72 ± 0.013 ¯ 63.178 ± 0.005 0.002 A -C-

+0.00 1599.06 -0.13 ¯ +0.000 62.955 -0.005 -A-

2.36/0.093 2.21/0.087

0.25 4 x R 0.13 0.010 4 x R 0.005 63 Chemical Conversion 3.30 Coat Per MIL-C-5541 Class 3 R 0.13 0.063 1.52/0.060 1.40/0.055 0.025 0.076/0.003 0.130/0.005 B 8.51 0.335 (15° 0' ± 0° 15')

5.84 ± 0.076

0.230 ± 0.003

1.27 2 x R 9.98 0.050 0.393

View B From Figure 5-31

Figure 5-32. 6306 PAF Detailed Dimensions (Download Figure) 5-27 Md mm 02772REU7 in

45° 30' CL Keyway IV 4 x ¯ 0.750 *

D ¯ 0.254/0.010 S A S ° 14 30'* (Thru holes for separation switch and/or lateral restraint device) 3 x 90 deg*

¯ 60.63 *

III I

*Used for Secondary * Latch System Only II View C-C (Looking Fwd)

See Figure 5-34 S/C

Separation A Plane

C C PAF 1604.72 ± 0.13 ¯ 63.178 ± 0.005

0.050/0.002 A

Chord Line

° ° +0 15' 30° 0' 60 0' -0° 10'

(1604.72) ¯ -C- (63.178)

+0.25 17.48 -0.00 0.38 0.13 +0.010 2 X R 0.688 -0.000 0.015 0.005 0.010 S C S CL Keyway +0.13 5.54 -0.00 0.218 +0.005 -0.000 View D

Figure 5-33. 6306 PAF Spacecraft Interface Dimensional Constraints (Download Figure) 5-28 Md 02773REU7 mm in. For Section Marked Area = 942 mm2 /1.46 in2 ±15% I = 420.400 mm4 /1.01 in4 ±15% Applicable Length L = 50.8/2.0

*Dimension To Be Met 1601.0 ¯ *If Secondary Latch 63.03 *System Is Used

3.56 0.140 60° 0' 1470 ¯ 57.872 1.27 R 0.050 L 5.84 ± 0.08 0.230 ± 0.003 18.11 17.48 0.713 0.688

0.08/0.003 B 0.64 0.025 22.35 15° 0' ±15' 0.88 9.52 0.375 1604.72 ± 0.13 ¯ 63 63.178 ± 0.005 -C- Chemical Conversion Coat Per MIL-C-5541 0.050/0.002 -A- Class 3 -C- View A From Figure 5-33

+0.13 1600.84 -0.00 ¯ +0.13 63.025 +0.005 4.19 -0.000 -0.000 +0.005 -A- 0.165 -0.000

0.838 1.91 0.686 1.78 2 x 0.033 0.075 0.027 0.070

0.26 0.13 2 x R 0.010 0.005 View B

Figure 5-34. 6306 PAF Spacecraft Interface Dimensional Constraints (Download Figure) 5-29 Md 02740REU7.1

120 3 Note: The capability is provided as a guide for spacecraft design and is 110 subject to verification by coupled loads analysis. The levels shown are for a 10-ft-diameter metal fairing 100 and are conservative for the 10-ft-diameter composite fairing. These levels will be updated in a future revision 90 With 9.5-ft-dia Fairing

80 2

70 With 10-ft-dia Fairing

60 Allowable Region CG Distance From Separation Plane (m) CG Distance From Separation Plane (in.)

40 1 5 67891011103 lb

3 4 5103 kg Spacecraft Weight

Figure 5-35. Capability of the 6306 PAF on 7920 Vehicle (Download Figure) 5-30 Md 03277REU7

19.05 Spacecraft ¯ 0.750 Separation Clamp Lateral Restraint Device and/or Switch Pad

Latch Pivot and Guard

Secondary Latch Clamp Retainer Assembly

Secondary Latch Retention Cable

Secondary Compression Spring Latch Linkage

PAF

Section A-A

Figure 5-36. 6306 PAF Secondary Latch (Download Figure) 5-31 Md 02775REU7

I mm in.

G 16.51 ¯ 0.65 Actuator 90 deg Apart (4 Places) G

1327.404 ¯ 1423.162 ±0.127 52.260 ¯ 56.030 ±0.005

0.051/0.002 D

II IV

Keyway Location

4 X 45°0'

III

609.6 24.00

Figure 5-37. 5624 PAF Detailed Assembly (Download Figure) 5-32 Md 02776REU7.1 mm in 1473.962 ¯ (56.030) +0.000 1396.619 Ð0.254 ¯ +0.000 54.965 Ð0.010 21.590 0.127/0.005 D 0.850 2.032 0.08 13.513 ° 45° 0 +0 30' (0.532) Ð0° 0' 3.048 3.226 ±0.76 0.120 0.127 ±0.003

0.025/0.001 0.56 3.048 0.12 20° 0' ± 0° 15' B

¯ 50.000

0.381/0.015 D

View A From Figure 5-37 +0.000 1417.701Ð0.127 ¯ +0.000 55.815 Ð0.005 2.362 ÐDÐ 0.093 2.210 0.087 1404.087 ¯ 55.279 Do Not Break Sharp Edges

Chemical Conversion 0.254 63 0.010 Coat per MIL-C-5541, 2 x R Class 3 0.127 0.005 1.524 0.762 0.060 2 x R 0.030 1.270 Separation Plane 0.050

View B 6.096 0.24

Figure 5-38. 5624 PAF Detailed Dimensions (Download Figure) 5-33 Md 02777REU7 mm in

Spacecraft 1327.404 ¯ 52.260

Retainer

Separation Plane

Clampband

Spring Actuator PAF

View G-G Rotated 45-deg CCW

Figure 5-39. 5624 PAF Clamp Assembly and Spring Actuator (Download Figure) 5-34 Md See Figure 5-41 02778REU7 mm in Spacecraft F

Separation Plane -A-

1423.162 ±0.076 0.05/0.002 D ¯ PAF C 56.030 ±0.003 C 0.05/0.002 D

38.10 II 4 x ¯ 1.50 Area for 170-lbf Separation Spring

¯ 0.254/0.010 M A C S

(Area extends from the separation See plane and forward 14.986/0.590) Figure 5-41 4 x 45° 0' E E

¯ 52.260

I III

IV +0.254 D 17.475Ð0.000 View C-C (Looking Fwd) 0.694+0.010 Ð0.000 +0.127 5.537 Ð0.000 0.218 +0.005 Ð0.000

Chord Line 30° 0' ±0° 30' 0.254 ±0.127 2 x R 0.010 ±0.005

° 60° 0'+0 15' Ð0° 0'

View D

Figure 5-40. 5624 PAF Spacecraft Interface Dimensional Constraints (Download Figure) 5-35 Md 02779REU7 mm 3.048 in. R 4.826 0.12 R For Section Marked 0.190 7.62 0.30 Area = 214.2 mm2 /0.332 in.2 ±15% I = 8741 mm 4 /0.021 in.4 ±15% Applicable Length = 25.4 mm/1.0 in. 25.4 20°0' ±0°15' 1.00 -A- 0.025/0.001

± G 3.226 0.076 0.127 ± 0.003 +0°30' °0' 45 -0°0'

1397.000 ± 0.254 ± +0.13 ¯ 55.000 0.010 Conversion 1.778 -0.00 Coat per MIL-C-5541, 63 Class 3 +0.005 0.070 -0.000

View F -C- From Figure 5-40 0.381 2 x R 2.286 R 0.015 0.090

-A- +0.130 0.254 -0.000 2 x +0.005 +0.130 0.010 -0.000 3.556 -0.000 +0.005 0.140 +0.127 -0.000 1418.844-0.000 ¯ +0.005 55.860-0.000 -D- View G 0.762 R 0.03

14.986 0.590 3.810 19.050 -A- 0.150 0.750 Separation Plane 22.098 0.870 38.10 Min 1.50 Section E-E From Figure 5-40

Figure 5-41. 5624 PAF Spacecraft Interface Dimensional Constraints (Download Figure) 5-36 Md 02748REU7

70 Note: For Clampband with 3900 lb Preload. The 5624 PAF is new, and the capabilities 60 of the different fairings will be provided in a future update.

20 CG Distance From Separation Plane (inches) 10

0 1000 2000 3000 4000 5000

Spacecraft Weight (lb) Figure 5-42. Capability of 5624 PAF on 7900 Vehicle (Download Figure) the separation clamp assembly interfaces that have a fit-check will be conducted with the spacecraft been qualified for the Space Transportation System using the flight PAF. This is typically conducted (STS). These clamp assemblies are listed in Table prior to shipment of the spacecraft from the manu- 5-3. facturing location to the launch site. MDA person- In addition, MDA is developing a 168-cm (66- nel will be available to conduct this activity. The fit- in.) diameter clamp assembly for the Delta III check verifies the flight interfaces (mechanical and vehicle that is Delta II-compatible; and MDA is electrical) and clearances of any attached hardware. studying a 119-cm (47-in.) clamp assembly. The spacecraft must include all flight hardware so

5.5 TEST FITTINGS AND FIT-CHECK that adequate access and clearance can be demon- POLICY strated. The spacecraft contractor will provide a A PAF test fitting can be provided to the payload support stand for the PAF and the bolts needed to contractor to assist in the conduct of the environ- secure the PAF to it. Specific detail requirements mental testing that is needed to ensure the flight for the fit-check will be provided by the Delta readiness of the spacecraft. This fitting would be Program Office. returned after the testing is completed. In addition, 5.6 ELECTRICAL DESIGN CRITERIA Table 5-3. Separation Clamp Assemblies Presented in the following paragraphs is a de- Approximate Max flight Spacecraft PAF diameter preload flange angle scription of the spacecraft/vehicle electrical inter- (cm/in.) (N/lb) (deg) 115/45 30,248/6800 15 face design constraints. The discussion includes 123/48 25,355/5700 20 blockhouse-to-spacecraft wiring, spacecraft um- 136/53 34,696/7800 20 M029-T013-4/4/96-3:52 PM bilical connectors, aerospace ground equipment 5-37 Md (AGE), the grounding system, and separation and shielded pairs (40 wires, No. 14 AWG), 2 switches. twisted and shielded triplets (6 wires, No. 1/0 AWG), 8 50-Ω coax cables and 6 fiber-optic cables. 5.6.1 Blockhouse-to-Spacecraft Wiring Space is available in the blockhouse for installa- Provisions are made for monitoring the space- tion of the ground support equipment (GSE) re- craft from the blockhouse after its mating to the quired for spacecraft checkout. The space allocated launch vehicle and until liftoff. Wiring is routed for the spacecraft GSE is described in Section 6 for from a payload console in the blockhouse through SLC-17 and Section 7 for SLC-2. There is also a second-stage umbilical connector, through fairing limited space in the umbilical J-box for a buffer wire harnesses, and to the spacecraft or PAF via amplifier or other data line conditioning modules lanyard-operated quick-disconnect connectors. required for data transfer to the blockhouse. The Provisions have also been made for controlling and space allocated in the J-box for this equipment has monitoring the spacecraft from the 1SLS Opera- dimensions of approximately 303 by 305 by 203 tions Building when it becomes operational. mm (12 by 12 by 8 in.) at SLC-17A and B and 381 For a typical vehicle, a second-stage umbilical by 330 by 229 mm (15 by 13 by 9 in.) at SLC-2. connector (JU2) is provided for payload servicing The standard interface method is as follows: wiring, of which 16 pins are reserved for vehicle A. The spacecraft agency normally provides a functions. A typical baseline wiring configuration console and a 12.2-m (40 ft) cable to interface with provides up to 31 wires through each of two fairing sectors. The fairing wire harnesses terminate in the spacecraft junction box in the blockhouse. The 32-pin lanyard disconnect connectors, which mate Delta program will provide the interfacing cable if to the PAF or directly to the spacecraft. Additional requested by the spacecraft agency. Once the 1SLS wiring can be provided by special modification. Operations Building comes on-line, the spacecraft Available wire types are twisted/shielded pairs, control console will be located in the spacecraft single- shielded, or unshielded single conductors. A control room. Cable interface lengths are still to be typical vehicle wire harness configuration is shown determined. in Figure 5-43. Other configurations can be B. The spacecraft apogee motor safe and arm accommodated. circuit (if applicable) must interconnect with the Twenty-four additional wires are available pad safety officer’s console (CCAS only). through the second-stage umbilical (JU1), which is C. A spacecraft-to-blockhouse wiring schematic shared with other second-stage system functions. is prepared for each mission from requirements The baseline wiring configuration between the provided by the spacecraft agency. Fixed Umbilical Tower (FUT) and the blockhouse D. To ensure proper design of the is the following. At CCAS the configuration at spacecraft-to-blockhouse wiring, the following in- SLC-17A and SLC-17B consists of 60 twisted and formation, which must comply with the above re- shielded pairs (120 wires, No. 14 AWG), 12 twisted quirements, shall be furnished by the spacecraft and shielded pairs (24 wires, No. 16 AWG), and 14 agency: twisted pairs (28 wires, No. 8 AWG). At VAFB the ■ Number of wires required. configuration at SLC-2 consists of 30 twisted and ■ Pin assignments in the spacecraft umbilical shielded pairs (60 wires, No. 12 AWG), 20 twisted connector(s). 5-38 Md H54663.1

P1118 P1103 J1103 JU2 P1115 P1100 J1100 JU2

AWG 20 A AWG 20 67 AWG 20 A AWG 20 132

B 76 B 120 C 77 C 121 D 57 D 150 E 43 E 151 F 44 F 152 AWG 20 G AWG 20 33 AWG 20 G AWG 20 161 AWG 16 H AWG 16 12 AWG 16 H AWG 16 176 AWG 16 J AWG 16 17 AWG 16 J AWG 16 181

AWG 20 K AWG 20 46 AWG 20 K AWG 20 153 L 65 L 130 M 66 M 131 AWG 20 N AWG 20 47 AWG 20 N AWG 20 154

AWG 16 P AWG 16 10 AWG 16 P AWG 16 186 AWG 20 R AWG 20 34 AWG 20 R AWG 20 140 S 54 S 141 T 55 T 163 U 35 U 164 V 36 V 165 AWG 20 W AWG 20 37 AWG 20 W AWG 20 180 AWG 16 X AWG 16 16 AWG 16 X AWG 16 185 AWG 16 Y AWG 16 21 AWG 16 Y AWG 16 170 AWG 20 Z AWG 20 24 AWG 20 Z AWG 20 170 *A 25 *A 171 *B 26 *B 172 *C 27 *C 173

*D 15 *D 178 179 *E 19 *E 182 *F 6 *F 183 *G 7 *G 188 AWG 20 *H AWG 20 8 AWG 20 *H AWG 20 189 *J 9 *J 187

Spacecraft/PAF (2 Stage) Spacecraft/PAF (2 Stage) 3rd Stage/Fairing Interface (3 Stage) 3rd Stage/Fairing Interface (3 Stage)

Delta II Payload Wiring Ð Quad I Delta II Payload Wiring Ð Quad III

Figure 5-43. Typical Delta II Wiring Configuration 5-39 Md ■ Function of each wire, including voltage, cur- Typical one-way line resistance for any wire is rent, frequency, load type, magnitude, polarity, and shown in Table 5-4. maximum resistance or voltage drop requirements. 5.6.2 Spacecraft Umbilical Connectors ■ Shield requirements for RF protection or signal noise rejection. For spacecraft configurations in which the um- ■ Voltage of the spacecraft battery and polarity of bilical connectors interface directly to the payload the battery ground. attach fitting, the following connectors (conform- ■ Part number and item number of the spacecraft ing to MIL-C-26482) are recommended: umbilical connector(s) (compliance required with ■ MS3424E61-50S (flange-mount receptacle) ■ the standardized spacecraft umbilical connectors MS3464E61-50S (jam nut-mount receptacle) listed in Section 5.6.2). These connectors mate to an MS3446E61-50P ■ Physical location of the spacecraft umbilical con- rack and panel mount interface connector on the nector including (1) angular location in relation to payload attach fitting. the quadrant system, (2) station location, and (3) For spacecraft configurations in which the um- radial distance of the outboard face of the connector bilical connectors interface directly with the fairing from the vehicle centerline for a fairing disconnect wire harnesses, the following connectors (conform- or connector centerline for PAF disconnect. ing to MIL-C-26482) are recommended: ■ ■ Periods (checkout or countdown) during which MS3470L18-32S (flange-mount receptacle) ■ hard-line controlled/monitored systems will be MS3474L18-32S (jam nut-mount receptacle) operated. These connectors mate to a 32-pin lanyard dis- During on-pad checkout, the spacecraft can be connect plug (MDC part number ST290G18N32PN) operated with the fairing installed or stored. Typi- in the fairing. cal harness arrangement for both configurations are The following alternative connectors, made by shown in Figure 5-44 for the ER and Figure 5-45 for Deutsch and conforming to MIL-C-81703, may be the WR. used when spacecraft umbilical connectors inter- Each wire in the baseline spacecraft- face with fairing-mounted wire harnesses or the to-blockhouse wiring configuration has a payload attach fitting: current-carrying capacity of 6 A, wire-to-wire iso- ■ D817*E61-OSN lation of 50 meg, and voltage rating of 600 VDC. ■ D817*E37-OSN

Table 5-4. One-Way Line Resistance Fairing on* Fairing off** Length Resistance Length Resistance Location Function No. of wires (m/ft) (ohm) (m/ft) (ohm) CCAS Data/control 60 348/1142 2.5 379/1244 3.7 CCAS Power 28 354/1160 1.3 385/1262 1.8 CCAS Data/control 24 354/1160 6.2 385/1262 7.3 VAFB Data/control 60 480/1576 3.7 511/1678 4.9 VAFB Data/control 40 480/1576 5.5 511/1678 6.6 VAFB Power 6 480/1576 0.9 511/1678 1.4 *Resistance values are for two parallel wires between the fixed umbilical tower and the blockhouse. **Resistance values include fairing extension cable resistance. M029, T014, 4/4/96, 3:52 PM 5-40 Md H54660.4 Cable Network Spacecraft Fairing Sector P708 P707 Fairing Sector PAF P1115 P1118

P1100 Motor P1103 Extension Extension Cables* Cables* J1100 Spin Table J1103

Second Stage JU2

* Extension Cables PU2 * Removed Prior to * Fairing Installation P3 P2 P1

J3A J2A J1A Umbilical Adapter J-Box

Umbilical Tower Spacecraft Terminal Room Interconnect Interface J-Box Distribution J-Box

Blockhouse Spacecraft Interface J-Box

Cables Provided by Spacecraft Contractor (40-ft Long)

Spacecraft Console

Figure 5-44. Typical Payload-to-Blockhouse Wiring Diagram for Three-Stage Missions at SLC-17

■ D817*E27-OSN tor shall be within five degrees of the fairing sector ■ D817*E19-OSN centerline. The face of the connector shall be within ■ D817*E12-OSN two degrees of being perpendicular to the centerline. ■ D817*E7-OSN A typical spacecraft umbilical connector is shown If “*” is 0, the receptacle is flange mounted; if 4, the in Figure 5-46. There should be no surrounding receptacle is jam-nut mounted. spacecraft intrusion within a 30-degree half-cone These connectors mate to a D817*E-series lan- angle separation clearance envelope at the mated yard disconnect plug in the fairing or rack and panel fairing umbilical connector (Figure 5-47). Pull forces plug on the PAF. The connector shell size numbers for the lanyard disconnect plugs are shown in Table (i.e., 37, 27, etc.) also correspond to the number of 5-5. For spacecraft umbilical connectors interfac- contacts. ing with the PAF the connector shall be installed so For spacecraft umbilical connectors that inter- that the polarizing key is oriented radially outward. face directly to the fairing wire harnesses, the space- Spring compression and pin retention forces for the craft connector shall be installed so the polarizing rack and panel connectors are shown in Table 5-6. key is in line with the longitudinal axis of the Separation forces for the bayonet-mate lanyard vehicle and facing forward (upward). The connec- disconnect connectors are shown in Table 5-7. 5-41 Md H54661.4 Cable Network Spacecraft Fairing Sector P708 P707 Fairing Sector PAF P1115 P1118

P1100 Motor P1103 Extension Extension Cables* Cables* J1100 Spin Table J1103

Second Stage JU2

* Extension Cables PU2 * Removed Prior to * Fairing Installation P1 P2 P3

J1 J2 J3 Spacecraft JBI Patch Panel J-Box

8 Delta JBI Patch Panel 6 Spacecraft Console 50½ coax.

62.5/125µÐm Cables Provided by Spacecraft Contractor Optic Fiber (40-ft Long) Spacecraft Blockhouse Equipment

Figure 5-45. Typical Payload-to-Blockhouse Wiring Diagram for Three-Stage Missions at SLC-2

Table 5-5. Disconnect Pull Forces (Lanyard Plugs) Table 5-6. Disconnect Forces Maximum (Rack and Panel Connectors) Minimum force engagement and Con- Maximum spring Maximum pin Con- for disengagement nector Shell compression retention nector Shell disengagement force type size (lb) (Kg) (lb) (Kg) type size (lb) (kg) (lb) (kg) D817X 61 77 34.93 68 30.84 MS347X 18 8.0 3.63 35.0 15.88 37 48 26.77 50 22.68 D817X 61 7.0 3.17 49.0 22.21 27 46 92.80 46 20.86 D817X 37 6.0 2.72 44.0 19.96 19 45 20.41 46 20.86 D817X 27 4.0 1.81 40.0 18.14 12 36 16.33 38 17.24 D817X 19 3.0 1.36 38.0 17.24 7 18 8.16 20 9.07 D817X 12 2.0 0.91 34.0 15.42 M029, T016, 4/4/96, 3:52 PM D817X 7 1.5 0.68 20.0 9.07 M029, T015, 4/5/96, 8:47 AM Program Office. This switch should be located to interface with the vehicle at the separation plane or 5.6.3 Spacecraft Separation Switch within 25.4 mm (1 in.) below it. A special pad will To monitor vehicle/spacecraft separation, a sepa- be provided on the vehicle side of the interface. The ration switch can be installed in the spacecraft. The design of the switch should provide for at least 6.4 configuration must be coordinated with the Delta mm (0.25 in.) over-travel in the mated condition. 5-42 Md H54687.1 H54664.2

Typical Spacecraft Umbilical Opening

Spacecraft Umbilical Umbilical Plug Connector

30 deg

Battery Flight Plug

Ordnance Arming Plug Disconnect Lanyard 30 deg

Fairing Umbilical Connector Spacecraft Separation Envelope

Figure 5-47. Spacecraft/Fairing Umbilical Clearance Figure 5-46. Typical Spacecraft Umbilical Connector Envelope

Table 5-7. Disconnect Forces (Bayonnet-Mate Lanyards) 5.6.4 Spacecraft Safe and Arm Circuit Min Max Connector Shell type size (lb) (kg) (lb) (kg) The spacecraft apogee motor safe and arm circuit ST290X 12 8 3.63 20 9.07 (if applicable) must interconnect with the pad safety 14 8 3.63 30 13.61 officer’s console in the blockhouse or Operations 16 8 3.63 30 13.61 18 8 3.63 35 15.88 Building, when operational. An interface diagram 20 8 3.63 35 15.88 for the spacecraft console and the pad safety 22 8 3.63 40 18.14 console is given in Figure 5-49 for the existing 24 8 3.63 40 18.14 blockhouse configuration and Figure 5-50 for the Typical spacecraft separation switch concepts are Operations Building configuration. Circuits for the shown in Figure 5-48. The switch located over the safe and arm (S&A) mechanism “arm permission” separation spring is the preferred concept. An alter- and the S&A talk-back lights are provided. This native for obtaining spacecraft separation indica- link is applicable at SLC-17 only and is not required tion is via the vehicle telemetry system. at SLC-2.

5-43 Md H54665.1 Preferred Configuration Alternative Configuration

Separation Switch

Separation Clamp AA Note: Switch centerline to be AA within 6.35 mm/0.25 in. of AA separation spring centerline

Figure 5-48. Typical Spacecraft Separation Switch and PAF Switch Pad

H54666.1 SP06E-12-10S (Provided by MDA)

+28 V C C 28 V 28 V

7.2 k½ 7.2 k½

Safe Arm

Spacecraft Ret A A Contractor Ret B B Blockhouse D D Console Arm E E Control F F G G 28 V Cable Length Approx 20 ft S/C Arm Permission Switch Pad Safety Console

Figure 5-49. Spacecraft/Pad Safety Console Interface for SLC-17

5-44 Md

H54662.3

(new or modified) or (new

Spacecraft Console Spacecraft

Range Comm Interface Comm Range

Key switch Arm to PSSC to Arm switch Key

Monitor Power Monitor

Arm Permission Status Permission Arm

Ground When Safe When Ground

Ground When Armed When Ground

Arm Power to PSSC to Power Arm

Granted

PSSC S/C Arm Permission Arm S/C PSSC

Terminal Room Terminal

Modified Spacecraft Rack Spacecraft Modified

(Launch Service Bldg) Service (Launch

Remote Site Remote

(800 Feet) (800

MDA Cables MDA

2 New 2

SP06E-12-10S

Blockhouse

MDA Fiber-Optic Interface Fiber-Optic MDA

F/O

MDA Fiber-Optic Interface Fiber-Optic MDA

SOB-Computer Room SOB-Computer

Cable

(125 feet) (125

New MDA New

MS3116P12-1OP

28 VDC 28

SOB-LCC

Switch

S/C Arm S/C

Permission

Pad Safety Supervisor’s Console Supervisor’s Safety Pad

Arm

Safe

28 VDC 28

Spacecraft Permission Spacecraft

Pin G Pin S&A S/C the Arms VDC 28 of presence The Ð Spacecraft to Command Permission Arm PSSC

Position Arm the in

Pin F Pin is Switch Key Console B/H S/C indicates VDC 28 of presence The Ð PSSC the to input Status Position Switch Key Safe/Arm

Pin E Pin On is Switch Power Arm Console B/H S/C indicates VDC 28 of presence The Ð PSSC the to input Switch Power Arming

Pin D Pin Granted Permission indicates VDC 28 of presence The Ð PSSC from Status Position Switch Permission Arm

Pin C Pin PSSC the to input Power Monitor VDC 28 Panel B/H manufacturer Spacecraft

Pin B Pin position Arm indicates Ground a of presence The Ð PSSC the to input Status Position Arm S&A Pin A Pin position Safe Indicates Ground a of presence The Ð PSSC the to input Status Position Safe S&A

5-45 Md Section 6 tics support. Requirements are described in docu- LAUNCH OPERATIONS AT ments prepared using the government’s universal EASTERN RANGE documentation system format. Formal submittal of This section presents a description of Delta these documents to the government agencies is Launch Vehicle Operations associated with Space made by MDA. MDA and the spacecraft agency Launch Complex 17 (SLC-17) at the Cape Canaveral generate the Program Requirements Documents Air Station, Florida. Delta II prelaunch processing (PRD). and spacecraft operations conducted prior to launch The organizations that support a launch are shown are presented. in Figure 6-1. A spacecraft coordinator from the MDA-CCAS launch team is assigned for each mis- 6.1 ORGANIZATIONS sion to assist the spacecraft team during the launch MDA operates the Delta launch system and main- campaign by helping to obtain safety approval of tains a team that provides launch services to NASA, the spacecraft test procedures and operations, inte- USAF, and commercial customers at CCAS. MDA grating the spacecraft operations into the launch provides the interface to the Department of Trans- vehicle activities, and serving as the interface be- portation (DOT) for the licensing and certification tween the spacecraft and test conductor in the block- needed to launch commercial spacecraft using the house during the countdown and launch. Delta II. MDA also has an established working relationship with Astrotech Space Operations. Astrotech owns and operates a processing facility 6.2 FACILITIES for commercial spacecraft in Titusville, Florida, in In addition to those facilities required for the support of Delta missions. Utilization of these fa- Delta II launch vehicle, specialized facilities are cilities and services is arranged by MDA for the provided for checkout and preparation of the space- customer. craft. Laboratories, clean rooms, receiving and ship- MDA interfaces with NASA at Kennedy Space ping areas, hazardous operations areas, offices, etc., Center (KSC) through the Payload Management are provided for use by spacecraft project personnel. and Operations Directorate. NASA designates a The commonly used spacecraft facilities at the Launch Site Support Manager who arranges all of eastern range are the following: the support requested from NASA for a launch from A. Spacecraft nonhazardous payload processing CCAS. MDA has an established interface with the facilities (PPF): USAF Space and Missile Center (USAF SMC) 1. Astrotech Space Operations provides Delta II Program Office and the 45th Space Wing Buildings 1 and 1A. Directorate of Plans. The USAF designates a Pro- 2. NASA provides Buildings AE and AM. gram Support Manager (PSM) to be a representa- B. Hazardous processing facilities (HPF): tive of the 45th Space Wing. The PSM serves as the 1. Astrotech Space Operations provides official interface for all support and services Building 2. requested. These services include range Commercial spacecraft will normally be pro- instrumentation, facilities/equipment operation and cessed through the Astrotech facilities. The other maintenance, as well as safety, security, and logis- facilities described in this section are controlled by 6-1 H54541

SPACECRAFT CUSTOMER ❏ Processes Spacecraft ❏ Defines Support Requirements

AIR FORCE QUALITY AIR FORCE MDC CCAS SAFETY ❏ Provides Quality Assurance ❏ Support for Launch Vehicle Processes Launch Vehicle ❏ Approves Procedures/Operations ❏ Ensures Spacecraft Support ❏ Requirements Are Satisfied AIR FORCE ❏ Interfaces With Government, AIR FORCE 1st SLS 45th Space Wing ■ Safety, NASA, and Air Force ❏ Owner of Launch Site ❏ Provides Base Support and 1st SLS ❏ Controls Government Launches ■ Range Services ❏ Adviser for Commercial Use of Government Facilities NASA KSC ❏ Provides Specific Base Support ■ Items

ASTROTECH ❏ Provides Off-Base Spacecraft ■ Facilities

Figure 6-1. Organizational Interfaces for Commercial Users

NASA and USAF and will be used only under these items and become familiar with such regula- special circumstances for commercial spacecraft. tions. These regulations include those imposed by The spacecraft agency must provide its own test NASA and USAF and FAA (refer to Section 9). equipment for spacecraft preparations, including telemetry receivers and telemetry ground stations. 6.2.1 Astrotech Space Operations Facilities Communications equipment, including some an- The Astrotech facility is located approximately tennas, is available as base equipment for voice and 5.6 km (3 mi) west of the Gate 3 entrance to KSC data transmissions. near the intersection of State Road 405 and State Transportation and handling of the spacecraft Road 407 in the Spaceport Industrial Park in and associated equipment are provided by MDA Titusville, Florida (Figures 6-2 and 6-3). This facil- from any of the local airports to the spacecraft ity includes 7,400 m2 (80,000 ft2) of industrial space processing facilities, and from there to the launch site. Equipment and personnel are also available for that is constructed on 15.2 hectares (37.5 acres) loading and unloading operations. Shipping con- of land. tainers and handling fixtures attached to the space- There are six major buildings on the site, as craft are provided by the spacecraft project. shown in Figure 6-4. Shipping and handling of hazardous materials A general description of each facility is given such as EEDs, radioactive sources, etc., must be in below. For details of door sizes, hook height, etc., a accordance with applicable regulations. It is the copy of the Astrotech Facility Accommodation responsibility of the spacecraft agency to identify Handbook is available. 6-2 AAAAAAAAAAAAAAAAAAAAAH54542.2 City of Titusville AAAAAAAAAAAAAAAAAAAAASpace Launch Complex 41

Indian Space Launch Complex 40

ToOrlando AAAAAAAARiver AAAAAAAAAAAAAA 50 Vehicle John F. Kennedy Assembly Visitor’s Space Center AAAAAAAAInformationAAAAABuildingAAAAAAAAA Center (VAB) Area AAAAAAAA405 AAAAAAAAAAAAACape Canaveral Air Station KSC Industrial 407 AAAAAAAAAAAAAAAAAAAAA Area Airport Space Launch AAAAAAAAAAAAASouth AAAAAAAA Complex 36A/B

ASTROTECH Kennedy Pkway

ToOrlando AAAAAAAAAAAAAAAAAAAAA Skid Strip 1

Interstate95 AAAAAAAAAAAAAAAAAAAAABanana River ExpresswayBee-Line AAAAAAAAAAAAAAAAAAAAA Space Launch Complex AAAAAAAAAAAAAAAAAAAAA17A/B 528 1 SLS AAAAAAAAAAAAAAAAAAAAAA1A Operations Building

AAAAAAAAAAAAAAAAAAAAACity of Cape Canaveral AAAAAAAACity of Cocoa AAAAAAAAAAAAA

Figure 6-2. Astrotech Site Location

H54543 State Road 405 Kennedy Space Center Main Gate and H54544 Guard Shack Equipment Entrance N North Chaffee Drive

Grissom Pkwy Non-Hazardous N Work Area

Future

State Road 407 Expansion Area White Road

Bldg 5 Space Executive Airport Bldg 1 Bldg 4 Chaffee Drive Bldg 1A Badge Building 2 Exchange Status Board Bldg 2 ASTROTECH State Road 528 Orlando

Bldg 3 Hazardous Work Area Addison Canal Bldg 6

Figure 6-3. Astrotech Complex Location Figure 6-4. Astrotech Building Locations 6-3 Building 1/1A, the Nonhazardous Processing floor coverings made of an electrostatic-dissipating Facility, is used for final assembly and checkout of (high-impedance) epoxy-based material. The the spacecraft. It houses spacecraft cleanroom high ground-level floor plan of Building 1/1A is shown bays, control rooms, and offices. Antennas mounted in Figure 6-5, and the upper-level floor plan is on the building provide line-of-sight communica- shown in Figure 6-6. tion with SLC-17 and Building AE at CCAS. Building 1. The airlock in Building 1 has a floor Building 2, the Hazardous Processing Facil- area measuring 9.1 m by 36.6 m (30 ft by 120 ft) and ity, houses three explosion-proof high bays for a clear vertical ceiling height of 7.0 m (23 ft). It hazardous operations, including liquid propellant provides environmentally controlled external ac- and solid rocket motor handling operations, spin cess to the three high bays and interconnects with balancing, third-stage preparations, and payload Building 1A. There is no overhead crane in the final assembly. airlock. Three RF antenna towers are located on the Building 3, the Environmental Storage Facil- roof of the airlock. The three high bays in Building ity, provides six secure, air-conditioned, 1 each have a floor area measuring 12.2 m by 18.3 masonry-constructed bays for storage of high-value m (40 ft by 60 ft) and a clear vertical ceiling height hardware or hazardous materials. of 13.2 m (43.5 ft). Each high bay has a 9072-kg Building 4, the Warehouse Storage Facility, (10-ton) overhead traveling bridge crane with a provides covered storage space for shipping con- maximum hook height of 11.3 m (37 ft). tainers, hoisting and handling equipment, and other There are two adjacent control rooms for each articles not requiring environmental control. high bay. Each control room has a floor area mea- Building 5, the Owner/Operator Office Area, suring 4.3 m by 9.1 m (14 ft by 30 ft) with a 2.7-m is an executive office building that provides the (8.9-ft) ceiling height. A large exterior door is spacecraft project officials with office space for provided in each control room to facilitate installa- conducting business during their stay at Astrotech tion and removal of equipment. Each control room and the Eastern Range. has a large window for viewing activities in the Building 6, the Fairing Support Facility, pro- high bay. vides covered storage space for launch vehicle Garment rooms provide personnel access to and hardware and equipment, and other articles not support the high bay areas. Limiting access to the requiring environmental control. high bays through these rooms helps control 6.2.1.1 Astrotech Building 1/1A. Building 1/ personnel traffic and maintains a cleanroom 1A has overall plan dimensions of approximately environment. 113 m by 34 m (370 ft by 110 ft) and a maximum Office accommodations for spacecraft project height of approximately 18 m (60 ft). Major fea- personnel are provided on the upper floor of Build- tures are two airlocks, four high bays with control ing 1 (Figure 6-6). This space is conveniently lo- rooms, and an office complex. The airlocks and cated near the spacecraft processing area and con- high bays are Class 100,000 cleanrooms, with the tains windows for viewing activities in the high bay. ability to achieve Class 10,000 or better cleanliness The remaining areas of Building 1 contain the levels using strict operational controls. They have Astrotech offices and shared support areas, includ- 6-4 Stair 2 Stair 1 H54545.4 131 125 119 130 124 118 115

114 116 133 129 128 127 123 122 121 117 111 112 132 113 Stair 2A Stair 1A 126 120 109 1103 101 1109 1121 Atrium 1108 110 102 1122 108 1123 103 1104 1117 134 135 136 AA 1111 1115 1124 1105 1113 1119 107 106 104 105 1118 1125

142 1106 1114 1116 1107 140 1112 137 141 1102 1101

Building 1A Building 1 1101 Large High Bay D 1113 Control Room D2 101 ASO Reception Area 122 Control Room B1 1102 Large Airlock 1114 Equipment Room 102 ASO Repro/Fax 123 Change Room B 1103 Mechanical Room 1115 Control Room D1 103 ASO Staff Office 124 Vestibule B 1104 Soundproof Conference 1116 Equipment Room 104 ASO Office Restroom 125 Storage B Room D1 1117 Office Area D1 105 ASO Staff Office 126 Restroom B 1105 Closet 1118 Break Room 106 ASO Staff Office 127 Control Room B2 1106 Restroom 1119 Corridor 107 ASO Staff Office 128 Control Room C1 1107 Restroom 1120 Not Used 108 Conference Room 129 Change Room C 1108 Vestibule 1121 Men’s Washroom 109 Women’s Restroom 130 Vestibule C 1109 Janitor Storage 1122 Men’s Restroom 110 Women’s Lounge 131 Storage C 1110 Not Used 1123 Janitor Closet 111 Men’s Restroom 132 Restroom C 1111 Change Room D 1124 Women’s Washroom 112 Break/Lunch Room 133 Control Room C2 1112 Air Shower 1125 Women’s Restroom 113 Janitor Closet 134 High Bay C 114 ASO Machine Shop 135 High Bay B 115 Corridor 136 High Bay A 116 Control Room A1 137 Common Airlock 117 Change Room A 138 Not Used 118 Vestibule A 139 Not Used 119 Storage A 140 Mechanical Room 120 Restroom A 141 Electrical Vault 121 Control Room A2 142 Telephone Room

Figure 6-5. First Level Floor Plan, Building 1/1A, Astrotech ing break room, supply/photocopy room, restroom and equipment is provided through a large exterior facilities, and a 24-person conference room. door. Building 1A. In addition to providing access via the The exterior wall of the airlock adjacent to the Building 1 airlock, Building 1A contains a separate exterior overhead door contains a 4.3-m by 4.3-m airlock that is an extension of the high bay and (14-ft by 14-ft) radio frequency (RF)-transparent provides environmentally controlled external ac- window, which looks out onto a far-field antenna cess. The airlock has a floor area measuring 12.2 m range that has a 30.5-m (100-ft) high target tower by 15.5 m (40 ft by 51 ft) and a clear vertical ceiling located approximately 91.4 m (300 ft) downrange. height of 18.3 m (60 ft). The airlock is a class The center of the window is 5.8 m (19 ft) above the 100,000 clean room. External access for payloads floor. 6-5 Stair 2 Stair 1 H54546.2 205 204

206 207 208 203 202 Stair 1A Stair 2A 201 2201 (134) 2204 2205 2206 (135) (136) 2215 2214 2213 2212 2211 2209 2207 2208 2203

2202 209 (1102) (1101) (137)

Building 1A Building 1 2201 Corridor 2209 Office Area D2 201 Telephone Room 2202 Corridor 2210 Not Used 202 Women's Restroom 2203 Break Room 2211 Office Area D3 203 Men's Restroom 2204 Men's Washroom 2212 Office Area D4 204 Janitor Closet 2205 Men's Restroom 2213 Office Area D5 205 Corridor 2206 Janitor Closet 2214 Conference Room D2 206 Office Area C 2207 Women's Washroom 2215 Office Area D6 207 Office Area B 2208 Women's restroom 208 Office Area A 209 Communications Room

Figure 6-6. Second-Level Floor Plan, Building 1/1A, Astrotech

The high bay has a floor area measuring 15.5 m and removal of equipment; and a large window for by 38.1 m (51ft by 125 ft) and a clear vertical ceiling viewing activities in the high bay. height of 18.3 m (60 ft). The high bay and airlock A garment room provides access for personnel share a common 27,215-kg (30-ton) overhead trav- and supports the high bay. Limiting access to the eling bridge crane with a maximum hook height of high bay through this room helps control personnel 15.2 m (50 ft). Personnel normally enter the high traffic and maintains a cleanroom environment. bay through the garment change room to maintain Office accommodations for spacecraft project cleanroom standards. The high bay is a class 100,000 personnel are provided on the ground floor and clean room. upper floor of Building 1A. This space is conve- There are two control rooms adjacent to the high niently located near the spacecraft processing area bay. Each control room has a floor area measuring and contains windows for viewing activities in the 9.1 m by 10.7 m (30 ft by 35 ft) with a 2.8-m (9.3- high bay. ft) ceiling height. Each control room has the follow- The remaining areas of Building 1A contain ing: a large interior door to permit the direct transfer shared support areas, including break rooms, of equipment between the high bay and the control restroom facilities, and two 24-person conference room; a large exterior door to facilitate installation rooms (one of which is a secure conference room 6-6 designed for the discussion and handling of classi- trench system is sloped so that any major spill of fied material). hazardous propellants would drain into the emer- 6.2.1.2 Astrotech Building 2. Building 2 has gency spill retention system. The center high bay overall plan dimensions of approximately 36.6 m contains an 8391-kg (18,500-lb) capacity dynamic by 36.6 m (120 ft by 120 ft) and a height of 14.0 m balance machine that is designed to balance solid (46 ft). Major features are two airlocks, three high rocket motor upper stages and spacecraft. bays, and two control rooms. The airlocks and high A control room is located next to each processing bays have floor coverings made of electrostatic- high bay to facilitate monitoring and control of dissipating (high-impedance) epoxy-based mate- hazardous operations. Visual contact with the high rial. They are Class 100,000 cleanrooms, with the bay is through an explosion-proof glass window ability to achieve Class 10,000 or better cleanliness in the separating wall. Access to the high bay area levels using strict operational controls. The ground- from the control room is via the garment room while level floor plan of Building 2 is shown in Figure spacecraft processing operations are being con- 6-7. ducted. The south airlock provides environmentally con- Because the spin balance table equipment in the trolled access to Building 2 via the south high bay. center high bay is below the floor level, other uses The south airlock has a floor area measuring 8.8 m can be made of this bay. The spin balance machine by 11.6 m (29 ft by 38 ft) and a clear vertical ceiling control room is separate from the spin room for height of 13.1 m (43 ft). There is no overhead crane safety considerations. Television cameras are used in the south airlock. for remote monitoring of spin room activities. The north airlock provides environmentally con- Adjacent to the south high bay, fuel and oxidizer trolled access to Building 2 via the north high bay cart storage rooms are provided with access doors to and, by restricting external Building 2 access to the the high bay and exterior doors for easy equipment south airlock, can be used as a fourth high bay. The access. Garment rooms provide personnel access to north airlock has a floor area measuring 12.2 m by and support the high bay areas. Limiting access to 15.2 m (40 ft by 50 ft) and a clear vertical ceiling the high bays through these rooms helps control height of 19.8 m (65 ft). The north airlock has a personnel traffic and maintains a cleanroom envi- 27,215-kg (30-ton) overhead traveling bridge crane ronment. with a maximum hook height of 16.8 m (55 ft). 6.2.1.3 Astrotech Building 3. The dimensions of Centered in the north airlock is a 7.6 m by 7.6 m Building 3 (Figure 6-8) are approximately 15.8 by (25 ft by 25 ft) floor area surrounded by a trench 21.6 m (52 by 71 ft). The building is divided into six system that drains into the emergency spill reten- storage bays, each with a clear vertical height of tion system. approximately 8.5 m (28 ft). The bays have indi- The north and south high bays are designed to support spacecraft solid propellant motor assembly vidual environmental control but are not cleanrooms, and liquid monopropellant and bipropellant trans- which mandates that payloads be stored in suitable fer operations. All liquid propellant transfer opera- containers. tions take place within a 7.6 m by 7.6 m (25 ft by 25 6.2.1.4 Astrotech Building 4. Building 4 (Figure ft) floor area surrounded by a trench system. The 6-9) is approximately 18.9 by 38.1 m (62 by 125 ft), 6-7 H54547

S N AAAAAAAAA E AAAAAAAAA AAAAAAAAA123 101 AAAAAAAAA AAAAA

102 103 104

121 AAAAAAAAA AAAA AAAAAAAAA AAAA 118 AAAAAAAAA AAAA AAAAAAAAA AAAA 119 106 105

116 115 114 113 111 109 117 108 122 112 107 110

101 South Airlock 113 Men's Restroom 102 South High Bay 114 South Change Room 103 Center High Bay 115 South Control Room 104 North High Bay 116 Balance Machine Control Room 105 Office 117 Mechanical Room 1 106 Mechanical Room 2 118 Corridor 107 Motor Generator Room 119 Oxidizer Cart Storage Room 108 North Control Room 120 Not Used 109 North Change Room 121 Fuel Cart Storage Room 110 Corridor 122 Electrical Vault 111 Women's Restroom 123 Building 2A Ð North Airlock High Bay 112 Janitor

Figure 6-7. Building 2 Detailed Floor Plan, Astrotech

6-8 H54548.1 vertical height which varies from 8.5 m (28 ft) along either sidewall to 9.7 m (32 ft) along the lengthwise

101 102 103 centerline of the room. While the storage area is protected from the outside weather, there is no environmental control.

108 The bonded storage area is environmentally con-

104 105 106 trolled and has a floor area measuring 3.6 by 9.7 m 107 (12 by 32 ft). 109 6.2.1.5 Astrotech Building 5. Building 5 (Figure N 6-10) provides office and conference rooms for the 101 Storage Bay A 102 Storage Bay B spacecraft project. 103 Storage Bay C 104 Storage Bay D 6.2.1.6 Astrotech Building 6. Building 6 (Figure 105 Storage Bay E 106 Storage Bay F 6-11) consists of a warehouse storage area and a 107 Panel Room 1 bonded storage area. The overall plan dimensions 108 Fire Equipment Room 109 Panel Room 2 of Building 6 are 15.2 m by 18.3 m (50 ft by 60 ft),

Figure 6-8. Building 3 Detailed Floor Plan, Astrotech with maximum roof height of 12.2 m (40 ft).

6.2.2 CCAS Operations and Facilities with a maximum roof height of approximately 9.1 Prelaunch operations and testing of Delta II space- m (30 ft). The major areas of Building 4 are the craft at CCAS take place in the following areas: warehouse storage area, bonded storage area, and A. Cape Canaveral industrial area. the Astrotech staff office area. B. SLC-17, Pad A or B. The large warehouse storage area has a floor area measuring 15.2 by 38.1 m (50 by 125 ft) and a clear 6.2.2.1 Cape Canaveral Industrial Area. A Delta II spacecraft facility located in the CCAS H54549.1 industrial area (Figures 6-12 and 6-13) is NASA- provided Building AE. This checkout facility is 106 105 103 102 complemented by the NASA-provided HPF, where 104 operations with liquid propellant, solid motor,

101 explosive device assembly, and dynamic balancing are conducted. A brief description of this facility follows. A copy of the Facility Handbook is available if addi- N tional details are required. 101 Warehouse 102 ASO Office In addition to this facility, several other USAF- 103 Bonded Storage 104 Restroom shared facilities or work areas at the CCAS are 105 Office Area A 106 Office Area B available for supporting spacecraft projects and the spacecraft contractors. These areas include the fol-

Figure 6-9. Building 4 Detailed Floor Plan, Astrotech lowing:

6-9 H54550

101 Lobby 102 Conference Room A 107 106 105 104 103 102 103 Office Area A 104 Office Area B 105 Office Area C 126 106 Office Area D 107 Office Area E 108 Office Area F 108 109 110 111 112 109 Office Area G 110 Office Area H 101 111 Office Area I 112 Office Area J 117 116 115 114 113 113 Mechanical Room 114 Office Area K 115 Office Area L 125 116 Office Area M 117 Office Area N 118 Office Area O 118 119 120 121 122 123 124 119 Office Area P 120 Office Area Q 121 Conference Room B 122 Kitchenette 123 Men’s Restroom 124 Women’s Restroom 125 Corridor 126 Corridor

Figure 6-10. Building 5 Detailed Floor Plan, Astrotech

H54581.2 ■ Solid-propellant storage area. N ■ Explosive storage magazines. Reference North ■ Electromechanical Test Facility. ■ Mission Director Center. ■ Liquid propellant storage area. 6.2.2.2 Building AE. Building AE (Figure 6-14), 101 Warehouse which is also called Hangar AE, is a Butler-type structure that is completely environmentally controlled. Located in the building are the Mission 102 Storage Director Center (MDC), Launch Vehicle Data Cen-

Room Function Length Width Height Doorway ter (LVDC), and offices for the spacecraft contrac- 101 Warehouse 60 (18.3) 50 (15.2) 40 (12.2) 20 x 40 tor and spacecraft management personnel. This (6.1 x 12.2) 102 Storage 20 (6.1) 10 (3.1) 8 (2.4) 3-0 x 6-8 building also houses the communications equip- (0.9 x 2.0) Notes: ment that links the Astrotech facility with NASA 1. All dimensions are approximate, and shown as feet (meters) and USAF voice and data networks at KSC and 1. or as feet-inches (meters). 2. The walls and ceilings in the warehouse are made of poly- CCAS. 1. covered insulation. The floor is made of concrete. Mission Director Center. Launch operations and Figure 6-11. Building 6 Detailed Floor Plan, Astrotech overall mission activities are monitored by the 6-10 (D3) H54551.4

Astrotech Mainland

Indian River

Vertical Assembly Complex 39 Kennedy Parkway Building (VAB) (Shuttle) KSC Industrial Area NASA Parkway

KSC Nuclear Fuel Storage SAEF 2 DMCO Bennett Causeway Solid Propellant Industrial Area Storage Area Banana River ❏ EMT

Area 55 Area 57

Atlantic Ocean

Cocoa Beach 1 SLS CCAS Operations Building Space Launch Complex 17 ❏ Pad A ❏ Blockhouse ❏ Pad B

Figure 6-12. Delta Spacecraft Checkout Facilities H54552.1

Building AE

E&O Building

SSC 112497 Figure 6-13. Cape Industrial Area 6-11 H54553 N

Main Entrance To Building AE

Launch Vehicle Data Center (LVDC)

NASA T/M Communications Mission Observation Ground Station Room Director Area Center

Figure 6-14. Building AE Floor Plan

Mission Director (MD) and the supporting mission of events after liftoff. These displays are used to management team in the Mission Director Center show present position and instantaneous impact (Figure 6-15) where the team is informed of launch point (IIP) plots. When compared with the theoreti- vehicle, spacecraft, and tracking network flight cal plots, these displays give an overall representa- readiness. Appropriate real-time prelaunch and tion of launch vehicle performance. launch data are displayed to provide a presentation 6.2.2.3 Spacecraft Assembly and Encapsula- of vehicle launch and flight progress. During launch tion Facility Number 2 (SAEF-2). The SAEF-2 is operations, the Mission Director Center also func- located at F Avenue and 7th Street in the Hypergol tions as an operational communications center from Maintenance Facility Area, KSC Industrial Area. which all communication emanates to tracking and The facility is used for the assembly, test, encapsu- control stations. Across the hall from the Mission lation, ordnance work, propellant loading, and pres- Director Center is the Launch Vehicle Data Center, surization of spacecraft. The facility contains ap- where MDA Delta management and technical sup- proximately 1556.08 m2 (16,750 ft2) of usable floor port personnel are stationed to provide assistance to space. Construction is of reinforced concrete and the launch team in SLC-17 blockhouse and the MD concrete block. The high bay is a steel frame struc- in the Mission Director Center. ture with insulated aluminum siding. At the front of the Mission Director Center are Like all launch support facilities assigned to a large illuminated displays that list the tracking particular project or resident contractor group, stations and range stations in use and the sequence SAEF-2 is available to other programs on a non- 6-12 AAAAAAAAAAAAAAAA H54609.1 AAAAAAAAAAAAAAAA AAAAAAAAAAAAAAAA AAAAAAAAAAAAAAAA1234567 891011 12 13 14 15 16 17 18 19 AAAAAAAAAAAAAAAA AAAAAAAAAAAAAAAA

AAAAAAAAAAAAAAAA20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 37

AAAAAAAAAAAAAAAAPAO AAAAAAAAAAAAAAAA AAAAAAAAAAAAAAAA AAAAAAAAAAAAAAAAAAAAAAAAAAAAAAAA38 39 40 41 42 43 44 45 46

AAAAAAAAAAAAAAAAObservation Room A AAAAAAAAAAAAAAAAAAAAAAAA AAAAAAAAAAAAAAAA

Figure 6-15. Building AE Mission Director Center interference basis if no other suitable facility is wide by 12.1-m-high (21-ft by 39.5-ft) horizontal available. sliding lift door separates the airlock from the Functional Divisions. Functionally, the building adjacent high bay. is divided into a clean work area (CWA) complex consisting of an airlock, a high bay, and two low 6.2.2.3.2 High Bay. The high bay measures 14.9 bays; a test cell; support office areas; and mechani- m by 30.2 m (49 ft wide by 99 ft long), providing a cal equipment rooms (Figure 6-16). Floors in the usable floor area of 450.7 m2 (4851 ft2); clear ceiling airlock, high bay, and test cell are designed for height is 22.6 m (74 ft) and is rated as a Class 3175.20 kg (7000 lb) per wheel plus 20% impact 100,000 CWA. Personnel and small equipment can loading. enter the high bay through the equipment airlock, 6.2.2.3.1 Airlock. The airlock, located at the north which is equipped with air showers. Clear access is end of the building, measures 12.5 m by 17.7 m (41 1.4 m by 2.1 m (4 ft 5 in wide by 6 ft 11 in high). ft wide by 58 ft long), providing a usable floor area 6.2.2.3.3 Low Bays. Two large bays are located of 221 m2 (2378 ft2) and is rated as a Class 300,000 along the west side of the high bay. One of the bays CWA. The airlock has a clear ceiling height of 15.9 measures 5.8 m wide by 21.9 m long (19 ft by 72 ft) m (52 ft). Access is gained by personnel doors, with a clear ceiling height of 7.62 m (25 ft); the other vestibule, and a 6.4-m by 12.2-m (21.5-ft wide by bay is 5.8 m wide by 8.2 m long (19 ft by 27 ft) and 40-ft high) vertical lift equipment door. A 6.5-m- has a clear ceiling height of 13.3 m (43.5 ft). The 6-13 H54608.4

Office Trailer NASA Office Trailer AA Office 121 Trailer AA123 (Room 201 Located Above Rooms 117 and 119) 101A 122 Storage

Airlock 119 118

101 Test Cell 117 Mech Equip AARoom Office Trailer 116 111 A AA115 102A 114 A 113 Change Room 112 Low Bay High Bay Clean Equip 109 AAAA Work Area Airlock 110 Area 108 Garment Storage and Issue A 107 Office 106 Entry Trailer 102 A 105 102B Mech Equip A Room 104 A Trench & A Drains Fuel/Oxidizer A Loading Area

103 Propellant N Loading Area

W E

Figure 6-16. SAEF-2 Floor Plan 6-14 combined bay areas provide a usable floor space of performed by spacecraft contractor personnel ac- 174.8 m2 (1881 ft2). The low bays are also rated as cording to range safety approved procedures. Class 100,000 CWAs. 6.2.4 Storage Magazines, CCAS 6.2.2.3.4 Test Cell. The test cell is located at the Storage magazines are concrete bunker-type struc- northeast corner of the facility. The cell measures 11.3 m by 11.3 m (37 ft by 37 ft), providing a usable tures located at the north end of the storage area. Only two of the magazines are used for spacecraft floor area of 127.2 m2 (1369 ft2 ). The clear ceiling height is 15.9 m (52 ft). Access to the test cell is ordnance. One magazine, designated MAG H, is ± ° gained through personnel doors and three 6.7-m- environmentally controlled to 23.9 2.8 C (75 ± ° wide by 12.2-m-high (22-ft by 40-ft ) vertical lift 5 F) with a maximum relative humidity of 65%. doors. The test cell can be used for spacecraft and This magazine contains small ordnance items such payload support activities not requiring Class as S&A devices, igniter assemblies, initiators, bolt 100,000 CWA conditions. cutters, electrical squibs, etc. 6.2.2.3.5 Other Rooms. The remainder of the The second magazine, designated MAG I, is used building consists of such support rooms as a me- for the storage of solid propellant motors. It is chanical equipment room, communication equip- environmentally controlled to 29.4 ± 2.8°C (85 ment room, entry and observation room, change ± 5°F) with a maximum relative humidity of 65%. room, storage room, miscellaneous equipment 6.2.5 Electrical-Mechanical Testing rooms, and limited office space. Facility, CCAS There is a 13.7-m-long by 8.5-m-wide (45-ft by The Electrical-Mechanical Testing Facility 28-ft ) conference area located on the second floor (EMT) (Figure 6-17), which is operated by range above rooms 117 and 199. contractor personnel, is used for such functions as Room 109 is used as the clean room garment ordnance item bridgewire resistance checks and storage and issue room. S&A device functional tests, as well as for test-firing 6.2.2.3.6 Trailers. There are five 3.7-m by 18.3- small self-contained ordnance items. m (12-ft by 60-ft) office trailers (Figure 6-16), Electrical cables that provide the interface available for payload personnel use when their between the ordnance items and the test equipment payload is in SAEF-2. Two are on the north side of already exist for most devices commonly used at the building; two on the west side; and one on the CCAS. These cables are tested before each use and east side. One of the trailers located on the north side the data is documented. In the event that a cable or is used by the NASA and PGOC facility managers. harness does not exist for a particular ordnance item it is the responsibility of the spacecraft project to 6.2.3 Solid Propellant Storage Area, CCAS provide the proper mating connector for the ord- The facilities and support equipment in this area nance item to be tested. A six-week lead time is are maintained and operated by the USAF range required for cable fabrication. contractor personnel. Ordnance item transport is The test consoles contain the items listed in Table also provided by range contractor personnel. Prepa- 6-1. The tests are conducted according to spacecraft ration of ordnance items for flight (i.e., S&A device contractor procedures which have been approved installation, thermal blanket installation, etc.) is by range safety personnel. 6-15 H54556.3 N Test Prep Chamber Bench

North Prep Prep Room Bench TV Camera

Work Ordnance Room TV Test Console Monitor Control Office TV Room Ordnance Monitor Test Console TV Monitor Control Lavatory

TV Camera Prep Bench South Prep Room

Test Prep Chamber Bench

Figure 6-17. Electrical-Mechanical Testing Facility Floor Plan

Table 6-1. Test Console Items towed to the pad with a MDA-provided tractor. The Resistance measurement Alinco bridge and null controls meter container (commonly called the handling can) can Digital current meter Resistance test selector be configured for either 2- or 3-stage missions. The Digital voltmeter Digital ammeter Auto-ranging digital voltmeter Digital stop watch height of the handling can varies according to the Digital multimeter Relay power supply High-current test controls Test power supply number of cylindrical sections required for a safe Power supply (5 V) Power control panel envelope around the spacecraft. The spacecraft High-current test power supply Blower M029-T018-4/26/93-9:46 AM container is purged with GN2 to reduce the relative humidity of the air inside the container and to 6.3 SPACECRAFT TRANSPORT TO LAUNCH SITE maintain a slight positive pressure. Temperature in the container is not controlled directly, but is main- After completion of spacecraft preparations and tained at acceptable levels when transporting the mating to the PAF in one of the Payload Processing spacecraft by selecting the time of day when move- Facilities (PPF) or Hazardous Processing Facilities ment occurs. The transportation environment is (HPF), the flight-configured spacecraft is moved to monitored with recording instrumentation. SLC-17 to join with the Delta II launch vehicle. MDA provides a mobile handling container to sup- 6.4 SLC-17, PADS A AND B (CCAS) port spacecraft transfer to the launch pad. SLC-17 is located in the southeastern section of The spacecraft handling can (Figure 6-18) is CCAS (Figure 6-12). It consists of two launch pads supported on a rubber-tired transporter and slowly (17A and 17B), a blockhouse, ready room, shops, 6-16 DAC 123087 H54344.3

Load Capacity (20,000 lb)

3657.6 144 6096 ~ 240

Extension Ladder Tool Box

4572 Wheel Base 180 3048 Track Width 120

m All dimensions are in in.

3048 3048 dia dia 120 120 Shackle (Inside Skin) Shackle (Inside Skin) Access Access Platform Platform Cover Cover

1171 AAAAA 1171 (Typical) (Typical) 46.12 AAAAA 46.12 AAAAA Payload (Reference) AAAAA Handling Can Handling Can (Shown with (Shown with 5 Cylindrical 4 Cylindrical AAAAA Sections) Sections)

Direct Mate Adapter for Handling Can 2-Stage Missions Configuration for 2-Stage Missions 1294 Conical Section for 3-Stage 50.93 Missions 6915 PAF (Ref) GSE Adapter Ring Clamp Handling Can Configuration for 3-Stage Missions AA AAAAA Direct Mate Adapter for 2-Stage Missions

Figure 6-18. Delta II Upper-Stage Assembly Ground-Handling Can and Transporter 6-17 and other facilities needed to prepare, service, and the blockhouse for the spacecraft consoles and launch the Delta II vehicle. console operations (Figure 6-25). After transition, The arrangement of SLC-17 is shown in Figure floorspace will continue to be allocated for remote 6-19 and an aerial view in Figure 6-20. controlled spacecraft consoles and battery charg- Since all operations in the launch complex area ing equipment. MDA will provide an interface for involve or are conducted in the vicinity of liquid or control and battery charging signals via fiber to the solid propellants and explosive ordnance devices, 1 SLS OB. Terminal board connections in the the number of personnel permitted in the area, spacecraft-to-blockhouse junction box (Figure 6- safety clothing to be worn, the type of activity 26 provide electrical connection to the spacecraft permitted, and equipment allowed are strictly regu- lated. Adherence to all safety regulations specified umbilical wires. in Section 9 is required. Safety briefings on these 6.4.3 First Space Launch Squadron subjects are given for those required to work in the Operations Building (1 SLS OB) launch complex area. In the mid-1997 to early-1998 timeframe, the 1 A changeout room is provided on the MST on SLS OB will be commissioned and all launch op- level 9 for use by spacecraft programs requiring this erations will be controlled from the Launch Control service. Center (LCC) located on the second floor of the 1 6.4.1 MST Spacecraft Work Levels SLS OB (Figure 6-27). The launch vehicle and GSE The number of personnel admitted to the MST is will be controlled and monitored from the OB via governed by safety requirements and by the limited the advanced launch vehicle control system (ALCS). amount of work space on the spacecraft levels. The There are two spacecraft control rooms and office relationship of the vehicle to the MST is shown in space, adjacent to the LCC on the second floor, Figure 6-21. Typical MST deck-level floor plans of available during processing and launch. Communi- pads 17A and 17B are shown in Figures 6-22, 6-23, cation equipment, located in each Control Room, and 6-24. will provide signal interface between the 1 SLS OB Outlets for electrical power and helium, nitro- and the blockhouse, as shown in Figure 6-28. Also, gen, and breathing air are provided on MST levels standard fiber optic communication bus interfaces 9A and 9B. See Figures 6-22 and 6-23 for locations. (i.e. EIA-422, RS-485, RS-488, EIA-232 will be Communications equipment provided on MST available for remote spacecraft equipment monitor- levels 9A, 9B, and 9C includes telephones and ing and control. operational communications stations for test sup- MDA is currently polling the spaecraft custom- port. See Figures 6-22, 6-23, and 6-24 for locations. ers to determine the best ways to meet their needs anticipating their interface requiremetns to include 6.4.2 SLC 17 Blockhouse PCM data, discretes, analog data, and RS232, 422, Launch operations will be controlled from the 485 and 488 IEEE standards. blockhouse until transition to the new 1st Space The range spacecraft interface spacecraft S&A Launch Squadron Operations Building (1 SLS OB). enable will be remoted from the PSSC console in Prior to transition to the 1 SLS OB, the blockhouse the OB to the ACS-R-B/H rack. The spacecraft is equipped with all of the vehicle monitoring and inteface will use the same pin connector definition control equipment. Floor space is also allocated in as the present spacecraft/PSSC Interface. 6-18 AA AA H54557.1

MST 17A MST 17B

Blockhouse

Delta Operations Support Building Delta Conference Lighthouse Road Room Trailer

N Horizontal Processing Building

Figure 6-19. SLC-17, CCAS

Figure 6-20. SLC-17 – Aerial View, Facing South 6-19 02879REU7

m All dimensions are in in. External Hoist 20,000 lb Capacity All station numbers are in inches Max Hook Ht = 171 ft, 10 in. Sta 47

Interior Bridge Crane 12,000 lb Capacity

Environmental Max Hook Ht = 163 ft Enclosure Sta 59

Elev 151 ft, 0 in. Sta 203.99 Elev 149 ft, 8-3/4 in. 10.0-ft dia Sta 219.22 Composite Fairing

3.63 143 3.25 9.5-ft dia Fairing 127.78

Level 9C Elev 139 ft, 1 in. Sta 347

3.05 Spacecraft Sta 414 120 3rd Stage

Level 9B Elev 129 ft, 1 in. Sta 467 3rd Stage Sta 500 2nd Stage 3.05 120 Sta 553.39

Level 9A Elev 119 ft, 1 in. Sta 588

3.28 129

Level 8B Elev 108 ft, 4 in. Sta 717

Figure 6-21. Environmental Enclosure Work Levels (Download Figure) 6-20 H54559 Downrange B C D EFG H 4.57 m 1.98 m 2.59 m 2.67 m 1.9 m 4.57 m 15 ft 0 in. 6 ft 6 in. 8 ft 6 in. 8 ft 9 in. 6 ft 3 in. 15 ft 0 in. AA 11 II 11 (2) Up Down (2)

AAHoist AA HOIST 4.27 m Fairing 14 ft 0 in. III 9 ft dia I AA Storage Area 25û Down Vestibule AA AA N 28û A Hoist IV AAAAAA AA To Cape (3) 11 Industrial AATV AA 3.66 m Area Hoist 12 ft 0 in. AC In AC In Airlock and Down A Changeout Room Pneumatic Panel Telephone Up (GN 2 , GHe, and Air)

m 0.6 & 2.4 Symbol Item Quantity Elevations Above Deck AC Inlets in. 24 & 96 Comm Box 6 Each 11 1.9 1.6 120 Vac 1φ 7 Outlets Hoists Comm Boxes 76 63 120/208 Vac 3φ 1 Outlet 1.86 1.7 Telephone 1 TV Cameras Pneumatic Panel 743 69

Figure 6-22. Level 9A Floor Plan, Pads 17A and 17B

H54560

Downrange BCDEFGH 4.57 m 1.98 m 2.59 m 2.67 m 1.9 m 4.57 M 15 ft 0 in. 6 ft 6 in. 8 ft 6 in. 8 ft 9 in. 6 ft 3 in. 15 ft 0 in.

AA II AAAA 11 (2) Down 4.27 m Hoist 11

AA Hoist Fairing AAAA 14 ft 0 in. (2) 12 ft dia Storage Area Up III 9 ft dia I AAAAAAAAA 25û With Inserts Crane Pendant Auxiliary Hoist Controls Vestibule AAAA N 28û Hoist IV A To Cape HoistAA AAAA 11 Industrial Area (3) Door Seal Controls AC In AC In Safety Belt Airlock 3.66 m Down 12 ft 0 in. Pneumatic Panel Telephone (GN 2 , GHe, and Air)

m Symbol Elevations Above Deck Item Quantity in. Comm Box 6 Each 2.4 1.8 & 3.9 11 φ AC Inlets 120 Vac 1 7 Outlets Hoists 96 72 & 152 φ 1 Outlet 120/208 Vac 3 1.3 1.6 1 Comm Boxes Telephone Pneumatic Panel 53 63

Figure 6-23. Level 9B Floor Plan, Pads 17A and 17B 6-21 H54561.2 Downrange CD E FG H 1.98 m 2.59 m 2.67 m 1.9 m 4.57 m 6 ft 6 in. 8 ft 6 in. 8 ft 9 in. 6 ft 3 in. 15 ft 0 in.

AA II A

AAHoist AAAA Hoist 12 ft dia III 11 ft dia I AAAAAAA With Inserts 6.10 m Dn 25û 20 ft 0 in. AAAA N 28û AAHoist IV AA AAAA To Cape Hoist Telephone TV Industrial Area AA AA AC In AC In

m Item Quantity Elevations Above Deck Symbol in. Comm Box 2 Each 2.4 Telephone 1 Hoists 96 1.8 & 3.9 AC Inlets 72 & 152 1.6 Comm Boxes 63

Figure 6-24. Level 9C Floor Plan, Pads 17A and 17B

H53562

Launch Conductor Launch Vehicle Control Consoles Spacecraft Coordinator

Pad Safety Console AA Spacecraft Area Adviser Consoles Spacecraft Blockhouse AA Consoles (Typical)

Spacecraft Blockhouse Telemetry Ground Station Junction Boxes

Figure 6-25. Blockhouse Console Locations 6-22 H54563.3

mm in. 914 36

(Cover Door Not Shown on Junction Box)

TB1 TB2 TB3 TB4 TB5

Crablock terminal blocks (PN A2S1415S) are provided by Delta for 12, 16, or 20 American Wire Gauge (AWG) 1067 wires. MDA will install the crablocks and 42 terminate the user's cable for the above AAA size wires 302 AAAAAA 8 Note: The distance from this terminal board to the spacecraft console area Delta Cables to Launch Area Access for Spacecraft Agency Cable is approximately 12.2 m (40 ft)

Figure 6-26. Spacecraft-to-Blockhouse Junction Box

H54610.1 Spacecraft Spacecraft Control Room Office No. 1 No. 1

Spacecraft Office and Library Computer and Control Room Facility UPS Mag Tape No. 2 Mech- Storage anical Civil Engineering Engineering Launch Support Area Women Control Center Engineering Systems Men Anomaly Room Facility Mechanical

Propellant UG QAM Test Engineering Chief CADD Conductor Engineer Facility Stairwell Electrical Mechanical Electrical Visitors Engineering Engineering Area and Comm Conference Room Room D-819683

Elev

Figure 6-27. Spacecraft Customer Accommodations – Launch Control Center 6-23

H54565.5

S/C UMB S/C

17B-GCR

17B-VCR2

17B-VCR1

17B-VCR2

17B-VCR1

17B

Terminal Room Terminal

17B 17B

ACS

Rack

ACS ACS

Rack

J-Box

Interface

S&A

Rack

J-Box

S/C-B S/C-B

S/C-B

Interface

ACS B/H ACS

B/H

ACS

TMR

Rack

S/C-B S/C-B

PSSC

Control

[FO/Cu]

Interface Interface

(485/FO)

(488/FO)

(422/FO)

(232/FO)

ACS-R-BH

(Analog/FO)

S/C interface S/C

(Discretes/FO)

Fiber Optics I/F Optics Fiber

SLC17 Blockhouse SLC17

B-CDP

A-CDP

F/O

144

Delta Delta

ACS

Work Work

Panels

Stations

Note: Cable length of path No. 1 is 20 feet 20 is 1 No. path of length Cable Note:

F/O

144 144

Delta Delta

[Cu/FO]

(485/FO)

(488/FO)

(422/FO)

(232/FO)

Interface Interface

ACS/PSSC ACS/PSSC

(Analog/FO)

S/C interface S/C

(Discretes/FO)

Fiber Optics I/F Optics Fiber

AAAAAAAAAAAAA

TMS

Rack

S/C-B S/C-B

PSSC

Control

ACS-R-OB

1 SLS OB SLS 1

TMR

B-CDP

A-CDP

ACS ACS

Work Work

Panels

Stations

Future - 1 SLS OB Control OB SLS 1 - Future

Current - Blockhouse Control Blockhouse - Current *Currently being defined being *Currently

Figure 6-28. Interface Overview - Spacecraft Control Rack in 1 SLS OB 6-24 6.5 SUPPORT SERVICES 6.5.1.2 Launch Decision Process. The launch decision process is conducted by the appropriate 6.5.1 LAUNCH SUPPORT management personnel representing the spacecraft, For countdown operations, the launch team is the launch vehicle, and the range. Figure 6-29 located in the blockhouse at SLC-17 and Hangar shows the typical communication flow required to AE with support from many other organizations. make the launch decision. For NASA missions, a The following paragraphs describe the organiza- Mission Director, launch management advisory tional interfaces and the launch decision process. team, engineering team, and quality assurance personnel will also participate in the launch decision 6.5.1.1 Mission Director Center (Hangar AE). process. The Mission Director Center provides the necessary seating, data display, and communication to 6.5.2 Weather Constraints control the launch process. Seating is provided for 6.5.2.1 Ground-Wind Constraints. The Delta II key personnel from MDA, eastern range, and the vehicle is enclosed in the MST until approximately spacecraft control team. For NASA launches, key L-7 hours. The tower protects the vehicle from NASA personnel also occupy space in the Mission ground winds. The winds are measured using an- Director Center. Government launches incorporate emometers at the 9.1-m (30-ft) and 28.0-m (92-ft) additional reporting and decision responsibility. levels of the tower.

H30486.2 Spacecraft Mission Director Center Spacecraft Ground Station (Hangar AE) Mission Control Center Spacecraft Spacecraft Spacecraft Network Spacecraft Spacecraft Spacecraft Status Network Status Status Voice Project Manager Mission Director Network (User) (User) Manager Spacecraft (User) Ground Station Spacecraft (User) Vehicle Status Launch Launch Spacecraft Director of Vehicle Mission Concurrence Status Director Mission Launch Engineering Director Advisory (USAF) Control Center (MDA) (MDA) Vehicle (User) System Status Launch Status Decision Launch Vehicle Systems Space Launch Range Operations Control Center Engineering Complex 17 Launch (MDA) Blockhouse Director (MDA) USAF (45 SW) Vehicle Status Status Status Launch Vehicle TOPS 1 Status Data Center (LVDC) Chief Field Launch Range Status Control (Hangar AE) Engineer Conductor Coordinator Office (MDA) (MDA) (MDA) (45 SW) ¥ Range Safety Status ¥ Eastern Range Status Spacecraft Status Site Coordinator Controller ¥ Weather (MDA) Status (USAF) ¥ Network Status Status

Figure 6-29. Launch Decision Flow for Commercial Missions–Eastern Range

6-25 The following limitations on ground winds (in- evaluated on launch day by conducting a trajectory cluding gusts) apply: analysis using the measured wind. A curvefit to the A. The MST shall not be moved from the Delta wind data provides load relief in the trajectory II if ground winds in any direction exceed 36 knots analyses. The curvefit and other load-relief param- (41 mph) at the 9.1-m (30-ft) level. eters are used to reset the mission constants just B. The maximum allowable ground winds at the prior to launch. 28.0-m (92-ft) level are shown on Figure 6-30 for 6.5.2.3 Weather Constraints. Weather con- 792X vehicles with lengthened nozzles on the air- straints are imposed by Range Safety to assure safe ignited GEMs. As noted on the figure, the con- passage of the Delta launch vehicle through the straints are a function of the predicted liftoff solid atmosphere. The following is a general overview of motor propellant bulk temperature. This figure ap- the constraints evaluated prior to liftoff. Appendix plies to both 9.5-ft and 10-ft diameter fairing con- B lists the detailed weather contraints. figurations. The plot combines liftoff controls, A. The launch will not take place if the normal liftoff loads, and on-stand structural ground wind flight path will carry the vehicle: restrictions. 1. Within 18.5 km (10 nmi) of a cumulo-

6.5.2.2 Winds Aloft Constraints. Measurements nimbus (thunderstorm) cloud, whether con- of winds aloft are taken at the launch pad. The Delta vective or in layers, where precipitation (or II controls and loads constraints for winds aloft are virga) is observed.

H54516.1 Delta II 7925/7925-10 Ground Wind Velocity Criteria N Six GEM Solids off the Pad, Three GEM LN Solids Air-Lit ER-Launch Pads 17A and 17B

Between 297¡ and 30¡ (92 ft) Temp (¡F) Wind Speed (knots) Vehicle Configuration: 792X Fairing Diameter: 9.5 & 10 ft 30 22 Solids: GEM LN 50 23 0 Launch Site: ER 330 30 Minimum Solid Motor Propellant Bulk 30¡Ð50¡F Temperature Range Anemometer Level: 92 ft

300 60 No Launch

Launch Angles Indicate Direction From Which Winds Come (Wind Speed Is Measured at 92 Feet) W 270 50 40 30 20 10 0 10 20 30 40 50 90 E Knots

Pad Azimuth 240 120 115 deg

35 knots (92 ft) Between 135¡ and 195¡ (92 ft) Temp (¡F) Wind Speed (knots) 150 210 30 25 180 35 26 S 40 27 45 28 50 29

Figure 6-30. Delta II 792X Ground Wind Velocity Criteria, SLC-17 6-26 2. Through any cloud, whether convective or B. First- and second-stage instrumentation may in layers, where precipitation or virga is be operated during an electrical storm. observed. C. If other vehicle electrical systems are pow- 3. Through any frontal or squall-line clouds ered when an electrical storm approaches, these which extend above 3048 m (10,000 ft). systems may remain powered. 4. Through cloud layers or through cumulus D. If an electrical storm passes through after a clouds where the freeze level is in the simulated flight test, all electrical systems are turned clouds. on in a quiescent state, and all data sources are 5. Through any cloud if a plus or minus evaluated for evidence of damage. This turn-on is 1 kV/m or greater level electric field con- done remotely (pad clear) if any category A ord- tour passes within 9.3 km (5 nmi) of the nance circuits are connected for flight. Ordnance launch site at any time within 15 min prior circuits are disconnected and safed prior to turn-on to liftoff. with personnel exposed to the vehicle. 6. Through previously electrified clouds not E. If data from the quiescent turn-on reveal equip- monitored by an electrical field mill net- ment discrepancies that can be attributed to the work if the dissipating state was short- electrical storm, a flight program requalification lived (less than 15 min after observed elec- test must be run subsequent to the storm and prior to trical activity). a launch attempt. B. The launch will not take place if there is precipitation over the launch site or along the flight 6.5.3 Operational Safety path. Safety requirements are covered in Section 9 of C. A weather observation aircraft is mandatory this manual. In addition, it is the operating policy at to augment meteorological capabilities for real- both MDA and Astrotech that all personnel will be time evaluation of local conditions unless a cloud- given safety orientation briefings prior to entrance free line of sight exists to the vehicle flight path. to hazardous areas such as SLC-17. These briefings Rawinsonde will not be used to determine cloud will be scheduled by the MDA spacecraft coordinator buildup. and presented by the appropriate safety personnel. D. Even though the above criteria are observed, 6.5.4 Security or forecast to be satisfied at the predicted launch time, the launch director may elect to delay the 6.5.4.1 Astrotech Security. Physical security at launch based on the instability of the current atmo- the Astrotech facilities is provided by chain link spheric conditions. perimeter fencing, door locks, and guards. Details 6.5.2.4 Lightning Activity. The following are of the spacecraft security requirements will be ar- procedures for test status during lightning activity: ranged through the MDA spacecraft coordinator. A. Evacuation of the MST and fixed umbilical 6.5.4.2 Launch Complex Security. SLC-17 tower (FUT) is accomplished at the direction of the physical security is ensured by perimeter fencing, Launch Conductor (Reference: Delta Launch Com- guards, and access badges. The MST white room is plex Safety Plan). a Defense Investigative Service-approved closed

6-27 area with cypher locks on entry-controlled doors. static proof test equipment. MDA’s operational Access can be controlled by a security guard on the team members are familiar with the payload pro- MST eighth level. cessing facilities at the CCAS, KSC, and Astrotech, 6.5.4.3 CCAS Security. For access to CCAS, US and can offer all of these skills and services to the citizens must provide full name with middle initial spacecraft project during the launch program. if applicable, social security number, company name, 6.6 DELTA II PLANS AND SCHEDULES and dates of arrival and expected departure to the MDA spacecraft coordinator or MDA/CCAS Secu- 6.6.1 Mission Plan rity. MDA Security will arrange for entry authority A mission plan (Figure 6-31) is developed for for commercial missions or individuals sponsored each launch campaign showing major tasks on a by MDA. Access by NASA personnel or NASA weekly timeline format. The plan includes launch sponsored Foreign nationals is coordinated by vehicle activities, prelaunch reviews, and space- NASA KSC with the USAF at CCAS. Access by craft PPF and HPF occupancy time. other US Government sponsored Foreign Nationals 6.6.2 Integrated Schedules is coordinated by their sponsor directly with the USAF at CCAS. For non-United States citizens, The schedule of spacecraft activities before inte- clearance information (name, nationality/citizen- grated activities in the HPF varies from mission to ship, data and place of birth, passport number and mission. The extent of spacecraft field testing var- date/place of issue, visa number and date of expira- ies and is determined by the spacecraft agency. tion, and title or job description) must be furnished Spacecraft/launch vehicle schedules are similar to MDA 2 weeks prior to the CCAS entry date. from mission to mission, from the time of space- Government sponsored individuals must follow craft weighing until launch. NASA or US Government guidelines as appropri- Daily schedules are prepared on hourly time ate. The spacecraft coordinator will furnish visitor lines for these integrated activities. These sched- identification documentation to the appropriate ules typically cover four days of integration effort agencies. After MDA Security gets clearance ap- in the HPF and eight days of launch countdown proval, entry to CCAS will be the same as for US activities. HPF tasks include spacecraft weighing, citizens. spacecraft third-stage mate and interface verification, and transportation can assembly around the 6.5.5 Field-Related Services combined payload. The countdown schedules pro- MDA employs certified propellant handler's en- vide a detailed, hour-by-hour breakdown of launch semble (PHE) suit propellant handlers, equipment pad operations, illustrating the flow of activities drivers, welders, riggers, and explosive ordnance from spacecraft erection through terminal count- handlers, in addition to people experienced in most down and reflecting inputs from the spacecraft electrical and mechanical assembly skills, such as project. These schedules comprise the integrating torquing, soldering, crimping, precision cleaning, document to ensure timely launch pad operations. and contamination control. MDA has under its Typical schedules of integrated activities from control a machine shop, metrology laboratory, LO2 spacecraft weighing in the HPF until launch (Fig- cleaning facility, proof-load facility, and hydro- ures 6-32 through 6-45) are shown as launch minus 6-28 H54567.2 APRIL MAY JUNE JULY AUGUST SEPT 20 27 4 11 18 25 1 8 15 22 29 6 13 20 27 3 10 17 24 31 7142128 DMCO C/O 17 PRESHIP MEETING - AT PUEBLO S.S. AT HPTF FIRST STAGE AT HPF PRE VOS MEETING - AT H/B 5 FIRST STAGE ERECTION 7 INTERSTAGE ERECTION VOS READINESS REVIEW - AT SLC 17 SOLID MOTOR ERECTION SOLID MOTOR PROCESSING SECOND STAGE ERECTION FAIRING ERECTION 13 S/C ARRIVE AT ASTROTECH PRETEST PREPS SPACECRAFT PROCESSING ASO BLDG 1 VEH SYS C/O SPACECRAFT PROCESSING IN ASO UPPER STG PROCESS ASO INTEGRATED OPERATIONS WEIGH SPACECRAFT 1 LAUNCH SITE READINESS REVIEW - AT SLC-17 4 PAYLOAD ERECTION 7 FLT PROG VERIF 8 ORDNANCE INSTAL 9 FAIRING INSTAL 10-11 FLIGHT READINESS REVIEW - AT SLC-17 11 MISSION REHEARSAL 14 S.S. PROP LOAD 14 LAUNCH READINESS REVIEW - AT SLC-17 15 RANGE SAFETY AND BEACON CHECKS 15 LAUNCH 16 Figure 6-31. Typical Mission Plan

H54568

0200 0300 0400 0500 0600 0700 0800 0900 1000 1100 1200 1300 1400 1500 1600 1700 1800 1900 2000 2100 2200 2300 2400

MORNING SCHEDULE BRIEFING AAAABAY OPENING CHECKS SETUP/CHECKOUT PWU AAAHOIST FUNCTIONAL/STRAY VOLTAGE AAPOSITION CLASS C WEIGHTS WEIGH SPACECRAFT ITEMS TO BE INSTALLED LATER AAAHYDRASET/LOAD CELL LINKAGE SETUP LOAD CELL SHUNT CHECKS AAAACLASS C WEIGH LIFT (VERIFY REPEATABILITY) * SPACECRAFT LIFT/WEIGH/LOWER * * SPACECRAFT LIFT/WEIGH/LOWER * SPACECRAFT LIFT/WEIGH/LOWER AAAAAAAAASPACECRAFT ACTIVITIES (REPEAT UNTIL TWO SUCCESSIVE TRIALS ARE WITHIN .02%) * NOTE: LIFT AND LOWERING STEPS TO BE SECURE LIFT EQUIPMENT AAAAAAAA ACCOMPLISHED BY S/C PERSONNEL.A AAASECURE WEIGH EQUIPMENT AAAAAAAAA AAAAABALLAST WEIGHTS (IF REQUIRED)

Figure 6-32. Typical Spacecraft Weighing (F-11 Day) 6-29 H54569

0200 0300 0400 0500 0600 0700 0800 0900 1000 1100 1200 1300 1400 1500 1600 1700 1800 1900 2000 2100 2200 2300 2400

VISHAY EQUIPMENT WARMUP MORNING SCHEDULE BRIEFING BAY OPENING CHECKS VISHAY/INSTRUMENTED STUD CALIBRATION

ACTUATOR INSTALLATION & LOCKWIRE

CLAMP BAND PREPARATIONS HOIST STRAY VOLTAGE CHECK LIFT/TRAVERSE/MATE SPACECRAFT AAAAAAASPACECRAFT ACTIVITIES SPACECRAFT TO PAF GAP MEASUREMENTS AAAAAAA CLAMP BAND INSTALLATION BAND TENSIONING/TAPPING SECURING

VISHAY RECHECKS S/C 3RD STAGE INTERFACE VERIFICATION (IF REQ'D)

Figure 6-33. Typical Mating of Spacecraft and Third-Stage (F-10 Day)

H54570

0200 0300 0400 0500 0600 0700 0800 0900 1000 1100 1200 1300 1400 1500 1600 1700 1800 1900 2000 2100 2200 2300 2400

MORNING SCHEDULE BRIEFING

BAY OPENING CHECKS

SEPARATION CLAMP BAND FINALIZING GAP MEASUREMENTS END FITTINGS INSTALL BAND RETAINERS

CONNECT SPRINGS TO RETAINERS CONNECT/TORQUE ETA INTO CUTTERS INSTALL ATTACH BOLT CUTTER BRACKETS LOCKWIRE SHIELDS/BRACKETS ETA

INSTALL NON-FLIGHT TAGS

AAAAAAA SEPARATION BLANKET INSTALLATION SPACECRAFT ACTIVITIES AAAAAAA FINAL INSPECTION PHOTOGRAPH ASSEMBLY

CLEAN & PREASSEMBLE CYLINDRICAL SECTIONS OF TRANSPORT CAN INSTALL/TORQUE 4 TRANSPORT CAN RING ASSEMBLIES TO SPIN TABLE

Figure 6-34. Typical Final Spacecraft Third-Stage Preparations (F-9 Day) 6-30 H54571

0200 0300 0400 0500 0600 0700 0800 0900 1000 1100 1200 1300 1400 1500 1600 1700 1800 1900 2000 2100 2200 2300 2400

MORNING SCHEDULE BRIEFING . BAY OPENING CHECKS ENGINEERING WALKDOWN CRANE FUNCTIONAL CHECKS CRANE STRAY VOLTAGE CHECKS HOIST INSPECTION EQUIPMENT PROOF LOAD VERIFICATION INSTALL CONICAL SHELLS INSTALL TEMP/HUMIDITY RECORDER INSTALL CYLINDER SHELLS BAG THE CAN ASSEMBLY AAAAAAASPACECRAFT ACTIVITIES REMOVE NOZZLE THROAT PLUG LIFT S/C & PAM/MATE TO TRAILER AAAAAAACOMPLETE PRIOR TO THIS DATE: TRAILER PURGE SETUP ❏ CLEAN/ASSEMBLE/TRANSPORT CAN PURGE CAN ASSEMBLY A❏ CLEAN/DISASSEMBLE/PREPAAAAA CANA ❏ FAB ILLUMALLOY BAG ATTACH IMPACT RECORDER ❏ CLEAN/MOVE TRAILER IN BAY A❏ 4 RINGAAAAA ASSY MATED TO SPINTABLEA AAAAAAA

Figure 6-35. Typical Installation of Transportation Can (F-8 Day)

H54572.1

2300 0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 TRANSPORTATION BRIEFING AT ASO XPORT S/C FROM ASO A/C WATCH & N204 /MMH MONITOR ERECTION PREPS PAD SAFETY SET UP HAZ. BADGE BOARD ERECTION BRIEFING UNDER HOOK ERECT & MATE SPACECRAFT BLOWDOWN & SAMPLE FAIRING DATA STATION WARMUP A/C DUCT PRIOR TO THIS DAY WHITEROOM STABILIZATION UNCAN SPACECRAFT AAAAAAASPACECRAFT ACTIVITIES INSTALL S/C SHROUD CONNECT S/C UMB CABLES INSTALL SPINTABLE BOLTS EMER POWER TEST BREAKDOWN & STOW S/C CAN F.S. PRELAUNCH HYD CKS CABLE UP THIRD STG BOATTAIL ENG'R W/D S.S. PRELAUNCH HYD CKS INITIATE S/C FUNCTIONAL CHECKS INITIATE S/C BATTERY CHARGE BATT PREPS & INSTL ODOTS RES CHECKS

Figure 6-36. Typical Spacecraft Erection (F-7 Day) 6-31 H54573.1

0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300 DATA STATION WARMUP RIFCA AIR ON PRETEST BRIEFING FOR FLIGHT PROGRAM VERIF TEST POWER ON & PRETEST PREPS AZIMUTH DETERMINATION PREPS S/C POWER ON & CONFIG TO LAUNCH MODE COMM CHECK MINUS COUNT (ABBRE.TERM.COUNT) AAAAAAASPACECRAFT ACTIVITIES SPACECRAFT POWER IN LAUNCH MODE T-0 AAAAAAA PLUS COUNT (FLT PRG VERIF TEST) ENGR. W/D & PARTIAL CENTER SECT CLS/OUT AZIMUTH DETERM. & MONUMENT CHECKS TEST RECYCLE & BATTERY CONNECT ENGR. W/D, PHOTOS & PART. GUID. SECT. CLOSEOUT S/C RECYCLE & CONFIGURE FOR STRAY VOLTAGE POWER ON STRAY VOLTAGE FILL LOX STORAGE TANK SAMPLE LOX STORAGE TANK A/C WATCH & N2H4 MONITOR

Figure 6-37. Typical Flight Program Verification and Stray Voltage Checks (F-6 Day)

H54574

0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300 BRIEFING DOTS T'SHOOT BRIEFING AAAAAAAA POWER OFF SV & ORD. CONN. SPACECRAFT ACTIVITIES REMOVE S/M NOSECONE DOORS AAAAAAAA S/M ENG'R W/D CENTER SECT ENG'R W/D 3RD STG, PAF & MINISKIRT ENG'R W/D INST'L 1/2 SEP COVERS DISCONNECT 3RD STG / FAIRING EXT CABLES SHEAR CORD INSTL (ROTATE S/C) REMOVE S/C SHROUD S.S. & PAF PREPS / CLEAN / INSP F.S. BOATTAIL CLOSEOUT & PREP FOR ETA HOOKUP CENTER SECT CLOSEOUT CLOSEOUT PHOTOS MST LVL CONFIG (LVLS 1A&6) A/C WATCH & N2H4 MONITOR AAAAAAS/C PRE-FAIRING CLOSEOUT

Figure 6-38. Typical Ordnance Installation and Hookup (F-5 Day) 6-32 H55179.3 0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300 FAIRING BAG REMOVAL HOIST FUNCTIONALS & HOIST BEAM / FAIRING CONNECTION S/C PREPS FOR FAIRING INSTL LEGEND S/C ENG WALKDOWN PAD OPEN SAMPLE FAIRING AIR AND STOW DUCT FAIRING INSTL BRIEFING AFLASHING AMBERÐ RAISE LVL 9B AND 9C LIMITED ACCESS INSTALL FAIRING SECTOR III, IN PEDESTAL (RESTRAIN AT 9C) AFLASHING REDÐ INSTALL FAIRING SECTOR II, IN PEDESTAL PAD CLOSED TERMINATE S/C PURGES LOWER NORTH LVL 9B, RESTRAIN AT 9B, INSTALLATION GUIDE DEVICES, REM 9C S/C ACTIVITY RESTRAINT, LOWER NORTH LVL 9C INSTALL RESTRAINT AND GUIDE DEVICES ON 9B INSTALL FAIRING SECTOR I IN PEDESTAL, LOWER LVL 9B AND 9C INSTALL HALO AND MATE FAIRING SECTORS INSTALL FIELD JOINT BOLTS C/D PREPS S/C BATTERY TRICKLE CHARGE (NO AIR CONDITIONING)

SUPPORT: DELIVER BAS TRAILERS (2) TO SLC-17 WILTECH (PARTICLE AND HC SAMPLES) SECURITY ESCORT DELIVER VS-1 AND VS-2 SCRUBBERS S/C VIDEO HOIST SUPPORT

TV AND COMM TECH PHOTO (S/C DOCUMENTARY) AREA CONDITIONS

Figure 6-39. Typical Fairing Installation (F-4 Day) H55178.2

0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300

INSTALL FIELD JOINT BOLTS (CONTINUED FROM F6 DAY) LEGEND

CLOCK FAIRING (IF REQUIRED) CLEANROOM GARMENTS PAD OPEN LVL 9A AND ABOVE (NOT 8B)AA RAISE LVL 9B AND 9C FLASHING AMBERÐ AALIMITED ACCESS REMOVE STRONGBACKS FLASHING REDÐ PAD CLOSED LOWER LVL 9B AND 9C AA S/C ACTIVITY INSTALL FAIRING A/C

RECONNECT S.S. DOTS CABLES

LANYARD AND CONN INSTALLATION

S/M INSUL INSTALLATION

A/C WATCH

SPACECRAFT BATTERY TRICKLE CHARGE

SUPPORT: PHOTOGRAPHER (DOCUMENTARY)

AREA CONDITIONS

Figure 6-40. Typical Fairing Installation (F-4 Day) 6-33 H54576

0100 0300 0500 0700 0900 1100 1300 1500 AAAAAAA1700 1900 2100 2300 FLIGHT READINESS REVIEW SPACECRAFT ACTIVITIES DATA STATION WARMUP AAAAAAA 3RD STAGE & S/C S&A FUNCTIONAL CHECKS & NCS CHECKS

PROP LOAD PREPS BRIEFING

S.S. SERV / BAS PREPS

SET UP 7630 ECOLYZERS

FAIRING FINALING (IF REQ)

PROP CHILL

S/C FUNCTIONALS

S/C BATTERY TRICKLE CHARGE

HYDRAZINE MONITOR AND A/C WATCH

Figure 6-41. Typical Propellant Loading Preparations (F-3 Day)

H54577

0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300

PROP CHILL DATA STATION WARMUP FAIRING ORD INSTL BRIEFING FINAL PROP SERVICE PREPS & FINAL BAS PREPS

OXIDIZER LOAD LUNCH BREAK SPACECRAFT ACTIVITIES S/M NOSECONE AAAAAAA CLOSEOUT FUEL LOAD AAAAAAABAGGIE INSP & ELECT CHECK (OFF PAD) 2ND STG PROP SECURE MISSION REHEARSAL PROP WATCH

A/C WATCH, HYDRAZINE MONITOR

Figure 6-42. Typical Second-Stage Propellant Loading (F-2 Day) 6-34 H54578

0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300

DATA STAT W/UP HEAT RP-1 BRIEFING PREPS LAUNCH READINESS REVIEW AZIMUTH PREPS AAAAAAASPACECRAFT ACTIVITIES 2ND STG TURNON 1ST STG TURNON AAAAAAA COMM CHECK PREPS FOR MST REMOVAL SLEW CHECKS BEACON & RANGE SAFETY CHECKS AZIMUTH UPDATE SECURING 2ND STG ENGINEERING WALKDOWN PROP SYS PRPS & LINE LVL RP-1 2ND STG THERMAL BLNKT INSTL S/C FINAL ENGINEERING WALKDOWN CLOSEOUTS CLASS A ORDN HOOKUP RED TAG INVENTORY S/C BATTERY TRICKLE CHARGE

AIR CONDITIONING WATCH, PROP WATCH, & HYDRAZINE MONITOR

Figure 6-43. Typical Beacon, Range Safety, and Class A Ordnance (F-1 Day)

H54702.1

0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300

A/C & PROP WATCH HEAT/RECIRCULATE RP-1 (AS REQ'D) CAMERA SETUP AAAAAAA BRIEFING (0815) ENGINEERING WALKDOWN SPACECRAFT ACTIVITIES AAAAAAA MST PREPS & MOVE / F.S. FINAL PREPS LANYARD TENSIONING PREP FOR SOLID MOTOR ARM MST REMOVAL & SECURING S/C TURN ON & CONFIG FOR LAUNCH SOLID MOTOR ARMING S/C RF LINK CHECK LAUNCH DECK / PAD SECURING PRESS S.S. HE & N2 TO 1100 PSI&HEX FILL BORESIGHT MST CAMERAS & TURN ON SEARCH LIGHTS LCE PREPS & HOLD FIRE CHECKS BUILT IN HOLD (60 MINS) TERMINAL COUNT S/C BATTERY TRICKLE CHARGING

Figure 6-44. Typical Delta Countdown (F-0 Day)

6-35 H54703.1

T-MINUS 150140 130 120 110 100 90 80 70 60 50 40 30 20 10 4 4 0

TERMINAL COUNTDOWN INITIATION & BRIEFING(1650) ALL PERSONNEL NOT INVOLVED IN TERMINAL COUNT CLEAR SLC-17 (SOUND WARNING HORN) PSS CLEAR BLAST DANGER AREA 1ST STAGE RATE GYRO TURN ON 60 1ST STG HE & N2 PRESS MIN 2ND STG TANK PRESS & HEX TOP OFF 10 MIN BUILT FIRST STAGE FUELING BUILT IN 2ND STG HE & N2 PRESS HOLD IN WEATHER BRIEFING AIR COND HIGH HEAT ON HOLD C-BAND BEACON CHKS AT AT TÐ150 LOX LOADING TÐ4 F.S. HYD ON & RATE GYRO CHKS MIN POWER & SWITCH VERIFICATIONS AUTO SLEWS SLEW EVALUATIONS TOP OFF HE & N2 F - 0 DAY LAUNCH WINDOW COMMAND CARRIER ON OPEN CLOSE DESTRUCT CHECKS LOCAL 19:40:00 19:58:00 STATUS CHECKS GMT 23:40:00 23:58:00 ARM SPACECRAFT S&A 18 MINUTES PRESSURIZE FUEL TANK ARM 3RD STG & DESTRUCT S&A SPACECRAFT INTERNAL (T-4) S/C BATTERY CHARGING SPACECRAFT LAUNCH READY (100 SEC) LAUNCH GMT 2100 2110 2120 2130 2140 2150 2200 2210 2220 2230 2240 2250 2300 2310 2320 2340 2326 2336

Figure 6-45. Typical Terminal Countdown (F-0 Day)

(F-) workdays. Saturdays, Sundays, and holidays F-10 Spacecraft is lifted and mated to the third are not scheduled workdays and therefore, are not stage; clampband is installed, and initial clampband F-days. The F-days, from spacecraft mate through tension is established. launch, are coordinated with each spacecraft agency F-9 Final preparations are made prior to can-up to optimize on-pad testing. All operations are for- for both spacecraft and third stage, and spacecraft/ third stage interface verification is done, if required. mally conducted and controlled using launch count- F-8 The payload handling can is assembled down documents. The schedule of spacecraft around the spacecraft/third stage; handling can activities during that time is controlled by the MDA transportation covers are installed; the can is placed Launch Operations Manager. Tasks involving the on its trailer; and the handling can purge is set up. spacecraft or tasks requiring that spacecraft person- F-7 Tasks include transportation to the launch nel be present are shaded for easy identification. site, erection, and mating of the spacecraft/upper Preparation for a typical three-stage mission from stage to the Delta II second stage in the MST CCAS is as follows; spacecraft and third stage cleanroom. Preparations are made for the launch checkout are completed before F-11 day. vehicle flight program verification test. F-11 Tasks include equipment verification, pre- F-6 The launch vehicle flight program verifica- cision weighing of spacecraft, and securing. tion test is performed, followed by the vehicle 6-36 power-on stray voltage test. Spacecraft systems to ings. Another set of launch vehicle-specific sched- be powered at liftoff are turned on during the flight ules are generated, on a daily timeline, which covers program verification test, and all data are monitored a two- or three-month period to show the complete for electromagnetic interference (EMI) and radio processing of each launch vehicle component. An frequency interference (RFI). Spacecraft systems individual schedule is made for DMCO, third-stage to be turned on at any time between F-5 day and HPF, and launch pad. spacecraft separation are turned on in support of the 6.6.4 Spacecraft Schedules vehicle power-on stray voltage test. Spacecraft support of these two vehicle system tests is critical to The spacecraft project team will supply sched- meet the scheduled launch date. ules to the MDA spacecraft coordinator who will F-5 The Delta II vehicle ordnance installation/ arrange support as required. connection, preparation for fairing installation, and spacecraft closeout operations are performed. 6.7 DELTA II MEETINGS AND REVIEWS F-4, 3 Spacecraft final preparations prior to fair- During launch preparation, various meetings and ing installation include Delta II upper stage close- reviews take place. Some of these will require out, preparations for second stage propellant servic- spacecraft customer input while others allow the ing, and fairing installation. customer to monitor the progress of the overall F-2 Second stage propellant is loaded. mission. The MDA spacecraft coordinator will en- F-1 Tasks include C-band beacon readout, and sure adequate spacecraft user participation. azimuth update, followed by the vehicle Class A ordnance connection, spacecraft ordnance arming, 6.7.1 Meetings and final fairing preparations for MST removal, Delta Status Meetings. Status meetings are gen- second-stage engine section closeout, and launch erally held twice a week at SLC-17. These meetings vehicle final preparations. include a review of the activities scheduled and F-0 Launch day preparations include gantry accomplished since the last meeting, a discussion of removal, final arming, terminal sequences, and problems and their solutions, and a general review launch. Spacecraft should be in launch configura- of the mission schedule and specific mission sched- tion immediately prior to T-4 minutes and standing ules. Spacecraft user representatives are encour- by for liftoff. The nominal hold and recycle point is aged to attend these meetings. T-4 minutes. Daily Schedule Meetings. Daily schedule meetings are held in the SLC-17 conference room to 6.6.3 Launch Vehicle Schedules provide the team members with their assignments One set of facility-oriented three-week sched- and to summarize the previous or current day’s ules is developed, on a daily timeline, to show accomplishments. These meetings are attended by processing of multiple launch vehicles through each the launch conductor, technicians, inspectors, engi- facility; i.e., for both launch pads, DMCO, Hangar neers, supervisors, and the spacecraft coordinator. M, solid motor area, and each of the three PPFs as Depending upon testing activities, these meetings required. These schedules are revised daily and are held at either the beginning or the end of the first reviewed at the twice-weekly Delta Status Meet- shift. 6-37 6.7.2 Prelaunch Review Process Launch Site Readiness Review (LSRR). This Periodic reviews are held to ensure that the space- review is held prior to erection and mate of the craft and launch vehicle are ready for launch. The upper stage and spacecraft. It includes an update of Mission Plan (Figure 6-31) shows the relationship the activities since the pre-VOS review and verifies of the review to the program assembly and test flow. the readiness of the launch vehicle, third stage, The following paragraphs discuss the Delta II launch facilities, and spacecraft for transfer of the readiness reviews. spacecraft to the pad. Postproduction Review. This meeting, conducted at Pueblo, Colorado, reviews the flight hardware at Flight Readiness Review (FRR). This review is the end of production and prior to shipment to an update of actuals since the Pre-VOS and is CCAS. conducted to determine that checkout has shown Mission Analysis Review. This review is held at that the launch vehicle and spacecraft are ready for Huntington Beach, California, approximately three countdown and launch. Upon completion of this months prior to launch, to review mission specific meeting, authorization to proceed with the loading drawings, studies, and analyses. of second stage propellants is given. This review Pre-Vehicle-On-Stand (Pre-VOS) Review. This also assesses the readiness of the range to support review is held at Huntington Beach subsequent to launch and provides a predicted weather status. the completion of Delta mission checkout (DMCO), Launch Readiness Review (LRR). This review and prior to erection of the vehicle on the launch is held on L-1 day and all agencies and contractors pad. It includes an update of the activities since the post-production review at Pueblo, the results of the are required to provide a ready-to-launch statement. DMCO processing, and any hardware history Upon completion of this meeting, an okay to enter changes. terminal countdown is given.

6-38 Section 7 port. Requirements satisfied by NASA and/or USAF LAUNCH OPERATIONS AT are described in the government’s Universal Docu- WESTERN RANGE ment System (UDS) format. MDA and the space- This section presents a description of Delta craft agency generate the Program Requirements Launch Vehicle Operations associated with Space Document (PRD). Formal submittal of these docu- Launch Complex 2 (SLC-2) at Vandenberg Air ments to the government agencies is arranged by Force Base, California. Prelaunch processing of the MDA. Delta II is presented, as well as a discussion of For commercial launches, MDA makes all the spacecraft processing and operations that are con- arrangements for the payload processing facilities ducted prior to launch day. and services. The organizations that support a launch from VAFB are shown in Figure 7-1. A spacecraft 7.1 ORGANIZATIONS coordinator from the MDA-VAFB launch team is As operator of the Delta launch system, MDA assigned to each mission to assist the spacecraft maintains an operations team at VAFB that pro- team during the launch campaign by arranging for vides launch services to NASA, commercial, and support of the spacecraft, assisting in obtaining USAF customers. This team is augmented by MDA safety approval of the spacecraft test procedures personnel from the Delta launch operations team and operations, integrating the spacecraft opera- from Space Launch Complex 17 at Cape Canaveral tions into the launch vehicle operations, and, during Air Station, Florida. MDA provides the interface to the countdown and launch, serving as the interface the DOT for licensing and certification to launch between the spacecraft and test conductor in the commercial spacecraft using the Delta II. blockhouse. NASA is responsible for the SLC-2 launch facilities at VAFB and operates spacecraft processing 7.2 FACILITIES facilities at VAFB that are used in support of Delta In addition to those facilities required for Delta II missions. The MDA interface with NASA at VAFB launch vehicle processing, specialized facilities are is through the VAFB KSC resident office. The provided for checkout and preparation of the space- interface with NASA at KSC Florida is through the craft. Laboratories, cleanrooms, receiving and ship- Program Management and Operations Directorate. ping areas, hazardous operations areas, offices, etc., NASA designates a Launch Site Support Manager are provided for spacecraft project personnel. (LSSM) who arranges all the support requested A map of VAFB is shown in Figure 7-2. from NASA for a launch from VAFB. MDA has The commonly used facilities at the western established an interface with the 30th Space Wing launch site for NASA or commercial spacecraft are Directorate of Plans; the Western Range has desig- the following: nated a Range Program Support Manager (PSM) to A. Spacecraft payload processing facilities (PPF): be a representative of the 30th Space Wing. The 1. NASA-provided building: 836. PSM serves as the official interface for all support 2. Astrotech Space Operations: 1032. and services requested. These services include range 3. California Commercial Spaceport, instrumentation, facilities/equipment operation and Inc.: Building 375 maintenance, safety, security, and logistics sup- 7-1 H54600.2 SPACECRAFT CUSTOMER ❏ Processes Spacecraft ❏ Defines Support Requirements

Government Quality NASA Safety NASA/USAF ❏ Approves Procedures/Operations ❏ Provide Quality Assurance ❏ Provides Site Safety Support Support for Site/LV MDA VAFB ❏ Processes Launch Vehicle NASA GSFC USAF 30th Space Wing ❏ Ensures Spacecraft Support ❏ Adviser for Commercial Use ❏ Provides Base Support and ❏ Requirements Are Satisfied of Government Property Range Services ❏ Interfaces With Government, ■ Safety, NASA, and USAF NASA KSC Air Force Safety ❏ Launch Site Facility Manager ❏ Approves Hazardous ❏ Controls Government Launches Procedures/Operations ❏ Adviser for Commercial Use of Government Facilities ❏ Provides Spacecraft Processing Facilities ❏ Provides Specific Base Support Items

ASTROTECH California Spaceport ❏ Provides Delta II Spacecraft ❏ Provides Spacecraft Processing ■ Processing Facilities ■ Facilities

Figure 7-1. Launch Base Organization at VAFB for Commercial Launches

B. Hazardous processing facilities (HPF): airfield), transportation to and from the payload 1. NASA-provided building: 1610. processing facilities and to the launch site will be 2. Astrotech Space Operations: 1032. provided by MDA or NASA, as appropriate. Equip- 3. California Commercial Spaceport, ment and personnel are also available for loading Inc.: Building 375 and unloading operations. It should be noted that While there are other spacecraft processing fa- the size of the shipping containers often dictates the cilities located on VAFB that are under USAF type of aircraft used for transportation to the launch control, commercial spacecraft will normally be site. The aircraft carrier should be consulted for the processed through the commercial facilities of type of freight unloading equipment that will be Astrotech Space Operations or the California Space- required at the western range. Shipping containers port. Government facilities for spacecraft process- and handling fixtures attached to the spacecraft are ing (USAF or NASA) can be used for commercial provided by the spacecraft project. spacecraft use only under special circumstances Shipping and handling of hazardous materials (use requires negotiations between MDA, the space- such as EEDs, radioactive sources, etc., must be in craft agency and the USAF or NASA). The space- accordance with applicable regulations. It is the craft agency must provide its own test equipment responsibility of the spacecraft agency to identify for spacecraft preparations, including telemetry re- these items and become familiar with such ceivers and telemetry ground stations. regulations. These regulations include those im- After arrival of the spacecraft and its associated posed by NASA, USAF, and FAA (refer to equipment at VAFB by road or by air (via the VAFB Section 9). 7-2 H54343.6

To Santa Maria RR Casmalia

135 Pacific N

Southern Rd Rancho El Umbra VAFB Golf Course Rd

San SR 1 El Rancho Antonio Saddle Club Gate Rd Purisima Point SLCÐ2 Cross Rd 13 th 135 Delta Launch St Complex SLC-2W VAFB East Housing Tangair Rd Airfield fie NASA Hazardous Air ld Santa Maria Gate Rd HQ Processing Facility Calif Blvd SR 1 Pine Canyon Astrotech Payload Space Operations Rd Vandenberg th Village 35 St Pacific Ocean Lompoc Gate

Ocean Beach Park Mission Hills

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Cre Rd and Solvang ek Ynez 246 Rd Lompoc SLCÐ3 Ridge Ocean Ave

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Facilities AFB Boundary

Scale In Feet 0 6,000 12,000

Figure 7-2. Vandenberg Air Force Base 7-3 7.2.1 NASA Facilities on South VAFB craft and associated equipment. This assembly room The spacecraft facilities are located in the NASA contains a class 100,000 horizontal laminar flow support area on South VAFB (Figure 7-3). The cleanroom, 10.4 m (34 ft) long by 6.6 m (21.5 ft) spacecraft support area is adjacent to State High- wide by 7.6 m (25 ft) high with temperature control way 246 on Clark Street and is accessible through of 15.6°-26.7°C ± 1.1° (60°-80°F ± 2°). The front of the SVAFB South Gate. The support area consists the cleanroom opens to allow free entry of the of the spacecraft laboratory (Building 836), NASA spacecraft and handling equipment. technical shops, NASA supply, and NASA engi- The cleanroom has crane access in the front-to- neering and operations building (Building 840). rear direction only; however, the crane cannot oper- 7.2.1.1 Spacecraft Laboratory. The Spacecraft ate over the entire length of the laboratory without Laboratory in Building 836 (Figure 7-4) is divided disassembly because its path is obstructed by the into work and laboratory areas and includes space- horizontal beam that serves as the cleanroom di- craft assembly areas, laboratory areas, cleanrooms, vider. Spacecraft Laboratory 1 will also support computer facility, office space, conference room, computer, telemetry, and checkout equipment in a and the telemetry station. separate room containing raised floors and an Spacecraft Laboratory 1 consists of a high bay under-floor power distribution system. This room 20.4 m (67 ft) long, 9.8 m (32 ft) wide, 9.1 m (30 ft) has an area of approximately 334 m2 (3600 ft2). 2 2 high and an adjoining 334 m (3600-ft ) support Temperature control in this area is 21.1°C ± 2.8° area. Personnel access doors and a sliding door 3.7 (70°F ±5°). by 3.7 m (12 by 12 ft) connect the two portions of this laboratory. The outside cargo entrance door to Spacecraft Laboratory 2 has a 527 m2 (5670 ft2) the spacecraft assembly room in Laboratory 1 is 6.1 work area. Access to this area from the high-bay m (20 ft) wide by 7.8 m (25 ft, 7 in.) high. A bridge service area is provided by a 3.7-by 5.2-m (12-by crane, with an 8.8-m (29-ft) hook height and a 4545- 17-ft) roll-up door. There are three electric over- kg (5-ton) capacity, is available for handling space- head cranes available: a fixed 909-kg (1-ton) hoist

H54605

To SLC2 State Highway 246 NASA Spacecraft Laboratories (Bldg 836) Lompoc AAAAAAAAAAAAAAAAAAAAAAAA AAAAAA AAAAA VAFB South Gate AAA Data Transfer Arguello Blvd Clark Street N Antenna NASA E&O (450 ft high) (Bldg 840)

Figure 7-3. Spacecraft Support Area 7-4

H30484.7 Mezz

Sliding Doors Second Floor

Heater AAEquip. Storage W M Room Room Second Mech Floor Storage Equip. AA Room

NASA S/C TM Mezz Lab 3 Office Office Storage

Comm TM A/C Room Room Office

S/C TM Clean- Offices LVDC Shop room

S/C TM Station Lab 1

Equip. Clean- Room room Storage S/C Lab 2 Conf Room M W

Sliding Doors

N Legend: Second Floor Above Second Floor This Area

Figure 7-4. Spacecraft Laboratory (Building 836) 7-5 with a 7-m (23-ft) hook height, and two 909-kg (1- munications panels for specific mission require- ton) monorail hoists with 5.5-m (18-ft) hook heights. ments. This provides the LVDC personnel with A horizontal laminar flow Class 100,000 cleanroom, technical communications to monitor and coordi- 9.1 by 5.2 by 5.2 m (30 by 17 by 17 ft), is located in nate both prelaunch and launch activities. Video this laboratory for spacecraft use. One end of the data display terminals in the LVDC are provided for cleanroom is open to allow access. display of range and launch vehicle technical infor- Spacecraft laboratory 3 has three work areas with mation. a total of 2323 m2 (25,000 ft2). This laboratory is The high bay is a 30.5- by 61-m (100- by 200-ft) permanently assigned to the NOAA Environmental area serviced by a 22,727-kg (25-ton) crane with a Monitoring Satellite Program. 7.6-m (25-ft) hook height. This area is ideal for The Launch Vehicle Data Center (LVDC) (Fig- handling heavy equipment and loading or unload- ure 7-5) is an area containing 24 consoles for the ing trucks. The high bay is heated and has 30.5-m MDA Delta management and technical support (100-ft) wide by 9.1-m (30-ft) high sliding doors on personnel. These positions are manned during count- both ends. down and launch to provide technical assistance to 7.2.1.2 NASA Engineering and Operations Fa- the launch team in the SLC-2 blockhouse and to the cility. The NASA Engineering and Operations Mission Director in the MDC in Building 840. facility in Building 840 (Figure 7-6) is located on These consoles have individually programmed com- SVAFB at the corner of Clark and Scarpino Streets.

H30482.2

C C C C 6 12 18 24 TV TV TV TV

C C C C 5 11 17 23 C C C C 4 10 16 22 TV TV TV TV

C C C C 3 9 15 21 C C C C 2 8 14 20 TV TV TV TV

C C C C 1 7 13 19

Entrance Entrance Hallway Weather Station Speaker Phone

Figure 7-5. Launch Vehicle Data Center, Building 836, Vandenberg AFB 7-6 H30437.1

Bldg 840 Floor Plan, Second Floor

Vault B209

Mission Director AAAA B207 B207 Center Room

Observation AAAA

Women AAAA

Men AAAA

AAAA AAAA Bldg 840 Floor Plan, First Floor AA Break Room AAAA AAAA A AAAAAA A AAAAAA A B107 AAA Conference Room Lobby

Women Boiler Men Room AAAAA AAAAA

Figure 7-6. NASA Building 840 7-7 It contains the NASA offices, NASA contractor tions. Two closed-circuit TV display monitors pro- offices, Mission Director Center, Observation vide display of preselected launch activities. Room, conference room, and other office space. An Observation Room, separated from the MDC The Mission Director Center (Figure 7-7) pro- by a glass partition, is used for authorized visitors. vides 24 communication consoles for use by the Speakers in the room monitor the communication mission director, spacecraft and launch vehicle channels used during the launch. representatives, experimenters, display controller, and communications operators. These consoles have 7.2.2 NASA Facilities on North Vandenberg individually programmed communications for spe- The NASA Hazardous Processing Facility (Build- cific mission requirements. This provides the MDA ing 1610) is located approximately 3.2 km (2 miles) personnel with technical communications to moni- east of SLC-2 and adjacent to Tangair Road (Figure tor and coordinate both prelaunch and postlaunch 7-8). This facility (Figure 7-9) provides capabilities activities. for the dynamic balancing of spacecraft, solid mo- Video data display terminals in the MDC are tors, and combinations thereof. It is also used for provided for display of range and vehicle technical fairing processing, solid-motor buildup, spacecraft information. A Readiness Board and a Events Dis- buildup, mating of spacecraft and solid motors, play Board provide range and launch vehicle/space- ordnance installation, and loading of hazardous craft status during countdown and launch opera- propellants. It houses the Schenk Treble dynamic

H54603

C C C CON 16 VDL CON 8 CON 24

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V V V VM VM 23 VM 29 VM 35 VM 22 VM28 VM 34 C C CON 7 C CON 23 VDL CON 15 VDL VDL STA 4 STA 7 STA 12

39 C VM VM C C CON 6 CON 14 CON 22

STA 3 STA 6 STA 11 09 VM C C CON 5 CON 13 C CON 21 VDL VDL VDL

38 STA 3 STA 6 STA 11 VM V V CV VM 21 VM 27 VM 33 VM 20 VM 26 VM 32 C C C 37 CON 4 CON 12 CON 20 VM VDL VDL VDL STA 2 STA 8 STA 10

08 C C C VM CON 3 CON 11 CON 19

STA 2 STA 8 STA 10

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C C VM CON 2 C CON 18 VDL CON 10 VDL VDL STA 1 STA 5 STA 9 V V V VM 19 VM 25 VM 31 VM 18 VM 24 VM 30

14 C C VM VM C CON 9 CON 17 VDL CON 1 VDL VDL STA 1 STA 5 STA 9 Observation Room (207)

Figure 7-7. Mission Director Center—Building 840 7-8 H30479.7

Hazardous Processing Facility Building 1610

Control Building 1605 Earth Barricade

Guard Station

Roadblock (Used To Control Access)

Pump N Station

0 25 50 100 Storage Tank Scale in Feet

Figure 7-8. Hazardous Processing Facility balancing machine and equipment for buildup, opening 5.2 m (17 ft) wide and 9.1 m (29 ft 9 in.) alignment, and balancing of the upper stage solid high. The facility is equipped for safe handling of propellant motors and spacecraft. Composite spin the hydrazine-type propellants used on many space balancing of the spacecraft/third stage combination vehicles for attitude control and supplemental pro- is not required. Facilities consist of the Hazardous pulsion. In the high bay, there is an overhead bridge Processing Facility (Bldg 1610), Control Room crane with two 4545-kg (5- ton) capacity hoists. (Bldg 1605), guard station, and fire pumping sta- The working hook height is 10.7 m (35 ft). A tion. Hazardous operations are conducted in the spreader beam is available which allows use of both Spin Test Building, which is separated from the 5-ton hoists to lift up to 10 tons. This beam reduces the available hook height by 3 ft 2 in. The Schenck Control Room by an earth revetment 4.6 m (15 ft) Treble spin balancing machine is in a pit in the floor high. The two buildings are 47.2 m (155 ft) apart. of Building 1610 with the machines mating at 7.2.2.1 Hazardous Processing Facility. The surface level with the floor. The facility is a Class Hazardous Processing Facility (Figure 7-9) is an 100,000 clean facility with positive pressure main- approved ordnance-handling facility and was contained in the room to minimize contamination from structed for dynamic balancing of spacecraft and the exterior atmosphere. The positive-pressure clean solid rocket motors. It is 17.7 m (58 ft) long, 10.4 m air is provided by the air circulation and condition- (34 ft) wide, and 13.7 m (45 ft) high with personnel ing systems located in a covered environmental access doors and a flight equipment entrance door equipment room at the rear of the building. Person- 7-9 H30483.3

Environmental Equipment Room

Spin Balance Machine

AAA Storage Room Access AAA Ladder AAA A AAAAAAAAAAA AAAAAAAAAA Crane AAAAAAAAAccess x

Platform Microwave x AAAAAAAA Equipment x Fuel xxxx AAAAAAAADrain Mechanical Entry Room Equipment Room AAAAAAAA Clean Room AAAAAAAA(High-Bay) AAAAAAAA Airlock Room

Fire Main

Figure 7-9. Spin Test Building (Building 1610) 7-10 nel gaining entry to the cleanroom from the entry Phase II expansion in 1996, this facility will include room must wear appropriate apparel and must pass an additional approximately 500 m2 (5,400 ft2) of through an airlock. The airlock room has an access industrial space. door to the exterior so that equipment can be moved With Phase II completion there will be three into the cleanroom. major buildings on the site, as shown in

7.2.2.2 Control Room Building. The Control Figure 7-11. Room Building (Figure 7-10) contains a control A brief description of each building is given room, an operations ready room, a fabrication room, below. For further details a copy of the Astrotech and a mechanical/electrical room. The control con- Facility Accommodation Handbook is available. sole for the dynamic balancing system is located Building 1032, the Payload Processing Facil- within the control room. Television monitors and a ity, houses two explosion-proof high bays and an two-way intercommunications system provide con- explosion-proof air lock/high bay for non hazard- tinuous audio and visual monitoring of operations ous and hazardous operations, and is used for final in the Spin Test Building. assembly and checkout of the spacecraft, liquid propellant and solid rocket motor handling opera- 7.2.3 Astrotech Space Operations Facilities tions, third-stage preparations, and payload final The Astrotech facility is located on 24.3 hectares assembly. The Astrotech facility will be on the (60 acres) of land at Vandenberg AFB approxi- Vandenberg fiber optics network, which will pro- mately 3.7 km (2 mi) south of the Delta II launch vide basewide communications capability. Anten- complex (SLC-2) along Tangair Road at Red Road na towers mounted on the building offer the option (Figures 7-2 and 7-11). The initial phase of this for line-of-sight RF communication with SLC-2. facility includes approximately 1,000 m2 (10,600 Building M1030, the Technical Support Build- ft2) of industrial space. With completion of the ing, provides administrative facilities for space-

H30476.4

Building 1605

Mechanical and Electrical Equipment

N Control Operations Room Ready Room

Toilet Fabrication Room

Figure 7-10. Control Room Building 1605 Floor Plan 7-11 H54657.1 these areas are made of an electrostatic dissipating Astrotech Space Operations, LP Vandenberg AFB, CA (high impedance) epoxy-based material. The

1032 ground-level floor plan of Building 1032 is shown Payload Processing in Figure 7-12. The airlock/high bay in Building Facility 1032 has a floor area measuring 12.2 by 18.3 m (40 by 60 ft) and a clear vertical ceiling height of 13.7 North m (45 ft). It provides environmentally controlled external access for large equipment entry into the high bays. The airlock/high bay contains a 9072 kg (10-ton) overhead crane with an 11.3 m (37-ft) hook M1031 height that serves both the airlock/high bay and the Payload Control Fence Tangair adjoining high bay. The two high bays in Building Fence Center Road 1032 differ in size. The smaller high bay, which Red Road adjoins and is an extension of the airlock/high bay,

M1030 M1028 has a floor area measuring 12.2 by 18.3 m (40 by 60 Tech Communications Support Support ft) and a clear vertical ceiling height of 13.7 m (45 Building Building ft). As noted above, the 10-ton overhead crane in the airlock/high bay also serves the smaller high bay. Figure 7-11. Astrotech Payload Processing Area The larger high bay, which adjoins the smaller high bay, has a floor area measuring 15.3 by 21.4 m (50 craft project officials with office space for conduct- by 70 ft) and a clear vertical ceiling height of 20 m ing business during their their stay at Astrotech and (65 ft). The larger high bay has a 27,200-kg (30-ton) the Western Launch Site. overhead traveling bridge crane with a maximum Building 1028, the Communications Support hook height of 16.8 m (55 ft). Building, contains the fiber termination unit which Each high bay has an adjacent control room with is the communications interface bewteen the floor areas as shown in Figure 7-2. A large exterior Astrotech facility and the Vandenberg communica- door is provided in each control room to facilitate tions network. This building also contains addi- installation and removal of equipment. Each con- tional administrative facilities for spacecraft project trol room has a large window for viewing activities officials. in the high bay. Garment rooms support the high 7.2.3.1 Astrotech Building 1032. Building 1032 bay areas, and provide personnel access to them. has overall plan dimensions of approximately 46 by Limiting access to the high bays through these 27 m (150 by 90 ft) and a maximum height of rooms helps control personnel traffic and maintains approximately 20 m (65 ft). Major features are an a cleanroom environment. airlock/high bay, two high bays with control rooms, There are three airlocks/low bays, two adjoining and three airlocks/low bays. The airlocks and high the smaller high bay and one adjoining the larger bays are Class 100,000 cleanrooms demonstrated high bay. Each airlock/low bay has a floor area capability, with the ability to achieve Class 10,000 measuring 6.1 by 6.1 m (20 by 20 ft) and a clear or better cleanliness levels. The floor coverings in vertical ceiling height of 3.7 m (12 ft). A large 7-12 H54649.3

101 111 Existing

112 101B 103 113 101 Change Room 1 105 104 101A 116 H2-compatible 102 Restroom 1 10-ton, 37-ft H/H Bridge 625 sq ft 110 102 103 Control Room 1 114 115 625 sq ft crane 104 Mechanical Room 1 20 x 40-ft 0/H door access 105 Switch Gear AAAAAAAAAA January 96 106 High Bay 1 107 107 Low Bay 1A 400 AAAAAAAAA AAAAAAA H -compatible 108 Low Bay 1B sq ft 106 117 2 109 Airlock 30-ton, 55-ft H/H bridge 110 Air Shower 1 AAAAAAAAA AAAAAAA crane 111 Entryway 108 20 x 50-ft 0/H door access 112 Office AA400 AAAAAAA AAAAAAA Hazardous fuel transfer 113 Change Room 2 sq ft 114 Restroom 2 AA 2400 sq ft 3500 sq ft 115 Air Shower 2 116 Control Room 2 117 High Bay 2 All clean rooms class 118 Low Bay 2 100,000 (Functional at less than 10,000) 119 Mechanical Room 2 AAAAAAA AA118 201 HVAC Room 400 119 201 AAAAAAA109 AAsq ft AAAAAAA 2400 sq ft Legend Roll-Up Door AA Containment Islands

Figure 7-12. Astrotech Payload Processing Facility-Building 1032 exterior door is provided in each airlock/low bay with the California Spaceport is located on South and a roll-up door is located in the wall between the Vandenberg adjacent to the Spaceport. This pro- airlock/low bay and the adjoining high bay to cessing facility is called the Integrated Processing provide environmentally controlled external ac- Facility (IPF) because both booster components cess for small equipment entry into the high bays. and payloads (satellite vehicles) can be processed in The airlocks/low bays are also suitable for staging the building at the same time. This facility, origi- of liquid propellants. nally built to process classified Space Shuttle pay- 7.2.3.2 Astrotech Building M1030. Building loads, is now a part of the Spaceport facilities. It is M1030 provides 223 m2 (2,400 ft2) of office and composed of two basic areas: the Processing Areas conference room space for the spacecraft project. and the Technical Support Areas. Figures 7-13 and 7-14 illustrate the two major areas: the Processing 7.2.3.3 Astrotech Building 1028. Building 1028 Areas located on the north side of the building and provides 111 m2 (1,200 ft2) of office area for the the Technical Support Areas on the south side. spacecraft project. The cross-sectional view of the IPF shown in Figure 7-14 illustrates the relationships between the 7.2.4 California Spaceport Facilities Technical Support Area and the Processing Area The California Spaceport is located on South Level numbers. Level numbers are defined in feet Vandenberg immediately south and adjacent to above the SLC-6 launch mount. Rooms on two SLC-6. The payload processing facility associated levels (89 and 101) provide office space and techni- 7-13 H54530.4

AIRLOCK TRANSFER HIGH BAY TOWER AREA

CELL CELL CELL 6902 PAYLOAD 1 2 3 PROCESSING ROOM

PHE ROOM

SECURITY OFFICE BREAK MENS CLEAN ROOM AIR SECURE ROOM ROOM LOCK CLEAN OPS FAX DRESSING AREA ROOM MGR 1495 SQ FT WOMENS ROOM N SECURE CONFERENCE CLEAN ROOM OFFICE MIC ROOM 988 SQ FT W E 1125 SQ FT ROOM

S SECURE OFFICE 152 SQ FT Figure 7-13. California Spaceport —Plan VIew of the IPF

cal support rooms ranging from 14 to 150m2 (150 to H54529.2 1620 ft2). These floors contain both “dirty” and clean elevators, clean dressing areas, tool cleaning areas, a PHE change room, dressing rooms, showers, break room, conference room, and rest rooms. An airlock on Level 89 separates the Technical Support Area from the Processing Areas. 7.2.4.1 Processing Areas. There are six major 75-TON TOP OF CELLS HOOK LEVEL 141 150' (MAX) processing areas within the IPF: LEVEL 129 1. Airlock LEVEL 119 TECH SUPPORT AREAS PAD DECK LEVEL 109 2. High Bay LEVEL 100 LEVEL 101 LEVEL 99 3. Three Payload Checkout Cells (PCC) LEVEL 89 LEVEL 89 LEVEL 79 LEVEL 69 4. Transfer Tower Area LOADING DOCK LEVEL 69 LEVEL 64 5. Fairing Storage and Assembly Area (FSAA) LEVEL 59 PROCESSING AREAS LEVEL 49 6. Miscellaneous Payload Processing N S Rooms (PPR) There are seven levels on the processing side; six of these can be seen in Figure 7-14. The seventh (Fairing Storage and Assembly Area) can be seen in Figure 7-15. The Airlock and the High Bay are on

Figure 7-14. California Spaceport—IPF Cross-Sectional Level 64. The Payload Checkout Cells floor and the View Transfer Tower Area are on Level 69. In addition to 7-14 H55405.2

Fairing Storage and Alternate Assembly Area Access (Future) (Future) Payload Checkout Cell No. 2

E W

VCF Access

Transfer Tower Area High Bay Airlock

Figure 7-15. California Spaceport—Cutaway View of the IPF (looking south) the cell floor at Level 69, there are six platform (30-ft by 147-ft) High Bay is serviced by a 75-ton levels in each of the three Processing Cells: 79, 89, bridge crane. The hook height in the High Bay is 99, 109, 119 and 129. There are Payload Processing 26.3 m (86 ft, 4 in). Access to the High Bay is Rooms on each level, providing a total of seven through the 7.3-m (24-ft) wide, 8.5-m (28-ft) door rooms similar to the Payload Processing Room from the Airlock. shown in Figure 7-13, for small payload processing The three class 100,000 clean, 10.7-m by 13.4-m or processing support. Access is provided to the (35-ft by 44-ft) Payload Checkout Cells (PCC) are Processing Area through the airlock on Level 89 of serviced by a 75-ton bridge crane with a 24.8 m (81 the Technical Support Area. ft, 4 in) hook height. Each cell also has 5-ton crane Figure 7-15 illustrates the IPF as viewed in cut- support with a hook height of 21.9 m (71 ft, 11 in). away looking south and shows the location of the Access to each cell is through doors from the High seventh area, the Fairing Storage and Assembly Bay with a total opening of 6.4 m (21 ft, 2 in). Area. This class 100,000 clean area provides the Tables 7-1 through 7-7 detail some of the capa- option for fairing storage and build-up prior to bilities in each of the processing areas. They define encapsulating the payload in the Transfer Tower constraints, customer-provided equipment, and tech- Area. nical capability summaries in nine categories: Space/ Access to the IPF is through the 7.3-m (24-ft) Access, Handling, Electrical, Liquids, Pneumatics, wide, 8.5-m (28-ft) high main door on the west side Environmental Control, Safety, Security and Com- of the Airlock. The 9.1-m by 30.5-m (30-ft by 100- munications. ft) class 100,000 clean Airlock has two 5-ton over- Some dimensions of the Processing Areas are head bridge cranes with a hook height of 10.8 m (35 summarized in Figure 7-16. Also shown are the ft, 5 in). The class 100,000 clean, 9.1-m by 44.8-m crane envelopes for the 5-ton cranes in the Airlock; 7-15 Table 7-1. Vehicle Checkout Facility Table 7-2. Vehicle Checkout Area Constraints User-provided support Constraints User provided support ■ No secure communications ■ GN and He pressure carts ■ ■ 2 No secure communications GN2 and He pressure carts ■ No installed propellant ■ Spill aspirators ■ No installed propellant ■ Spill aspirators loading system loading system ■ No installed propellant ■ No installed propellant disposal system disposal system ■ No installed microwave or ■ No installed microwave or infrared intrusion detection infrared intrusion detection system system Capability type Capability Capability type Capability 1. Space/access ■ Floor loading for mobile equipment 1. Space/ ■ Floor loading for mobile equipment meets meets [AASHTO H-20] access [AASHTO H-20] ■ 30-by-100 ft internal floor space ■ Work space approximately 30 in by 146 ft 9 ■ 28-ft-ht by 24-ft-w door openings in ■ Adjacent to washdown area outside ■ Adjacent to Transfer Tower Area and ■ Accept tow vehicle/transporter of 20 Payload Checkout Cells by 90 ft 2. Handling ■ 75-ton overhead bridge main crane 2. Handling ■ Two 5-ton overhead bridge cranes (currently proofloaded to 29 tons) ■ Crane maximum hook height of 35 ft ■ Hook height 5 in Ð VCA 86 ft 4 in above floor (floor ■ Speeds at elev 64 ft) Ð Hoist 16 fpm Ð Checkout 81 ft 4 in maximum above Ð Bridge 14 fpm cells loor (floor at elev 69 ft) Ð Trolley 14 fpm Ð Transfer 81 ft 4 in maximum above ■ Pendant control at elevation 64 ft tower floor (floor at elev 69 ft) (floor) ■ Speeds 3. Electrical ■ Utility and technical power 120/208 Ð Hoist 10 fpm VAC Ð Bridge E/W 15 fpm and 30 fpm ■ Hazard-proof electrical equipment Ð Trolley N/S 15 fpm and 10 fpm as defined in the National Electrical ■ Micro drive Code Articles 500Ð516 Ð Hoist 0.5 and 1.5 fpm ■ Multi-point grounding per MIL-STD- Ð Bridge 0.5 fpm 1542 Ð Trolley 0.5 fpm 4. Liquids ■ Cleaning water supply ■ Two portable push-button stations with 60- Ð 100 gpm at 80 psig foot flex cable Ð 1.5-in. male hose thread ■ Connected to junction boxes on north wall 5. Pneumatics ■ Compressed air 125 psig 3. Electrical ■ Utility and technical power 120/208 VAC Ð 3/8-in. QD interface ■ Hazard-proof electrical equipment as 6. Environment ■ Buffer for operations between defined in the National Electrical Code external environment and High Bay Articles 500Ð516 ■ 100,000 clean room capability ■ Multi-point grounding per MIL-STD-1542 Ð Inlet air Class 5000 4. Liquids ■ None Ð Temp 70° ± 5° F 5. Pneumatics ■ Gaseous Nitrogen (GN2) Ð RH 30Ð50% 6. Environment ■ 100,000 clean room capability Ð Dif 0.05-in. Wg Ð Inlet air Class 5000 Ð Air chg 10Ð12 changes/hr min Ð Temp 70° ± 5° F ■ Central vacuum system Ð RH 30Ð50% ■ 7. Safety All electrical equipment is hazard- Ð Dif 0.05-in. Wg proof as defined in the National Ð Air chg 10Ð12 changes/hr min Electrical Code Articles 500Ð516 ■ Central vacuum system ■ Fire detection and suppression 7. Safety ■ All electrical equipment is hazard-proof as system defined in the National Electrical Code ■ 8. Security Access control Articles 500Ð516 Ð KeyCard/cipher system ■ Fire detection and suppression system Ð Intrusion detection system (BMS (suppression system currently inactivated) switches) 8. Security ■ Access control Ð Vault doors with S&G 3-position Ð KeyCard/cipher system tumbler Ð Intrusion detection system (BMS Ð Lockable personnel and switches) hardware access doors Ð Lockable personnel and hardware ■ 9. Communications Administrative phone access doors ■ Operational Voice System (OVS) 9. Communi- ■ Administrative phone ■ Area warning system cations ■ Operational Voice System (OVS) ■ Paging system ■ t2 Area warning system ■ Paging system t4 7-16 Table 7-3. Payload Checkout Cells Capabilities Constraints User-provided support ■ ■ No installed GN2 or GHe systems GN2 and He pressure carts ■ Hypergolic propellant loading system in Cell 2 not activated ■ Spill aspirators ■ No installed propellant disposal system Capability type Capability 1. Space/access ■ Design floor loading Ð 100 psf on checkout cell floor Ð 75 psf plus a 4000-lb load on four casters (4 by 6 ft) on fixed platforms Ð 50 psf plus a 1200-lb load on folding platforms ■ Work space approximately 35 by 44 ft ■ Cell door opening 21 ft 2 in x 71 ft high ■ Adjacent to Transfer Tower Area and High Bay ■ Six working platform levels (fixed and fold-down plus finger planks in Cells 2 and 3), spaced ten feet apart 2. Handling ■ 5-ton overhead bridge crane ■ Hook height (floor at elev 69 ft) Ð Cell 1 71 ft 6 in above floor Ð Cell 2 71 ft 11 in above floor Ð Cell 3 71 ft 4.5 in above floor ■ Speeds Ð Hoist 16 fpm (Cells 2/3) 10 fpm (Cell 1) Ð Bridge E/W 10 fpm Ð Trolley N/S 10 fpm (Cell 1) 5 fpm (Cell 2) 17 fpm (Cell 3) ■ Micro drive Ð Hoist 0.5 fpm Ð Bridge 0.5 fpm Ð Trolley 0.5 fpm ■ Portable push-button station with 45-foot flex cable connected to receptacle on northeast corner of cell on any level 3. Electrical ■ Utility and technical power 120/208, 408 VAC ■ Multi-point grounding per MIL-STD-1542 4. Liquids ■ Cleaning water supply Ð 50 gpm at 80 psig Ð 1-in hose bib with 1-in male hose thread on south wall of each level ■ Hypergolic 5. Pneumatics ■ Compressed air 125 psig 3/8-in. QD at two locations per cell 6. Environment ■ 100,000 clean room capability (Class 5000 HEPA) Ð Inlet air Class 5000 Ð Temp 55° to 75° ± 5°F selectable Ð RH 30Ð50% Ð Dif 0.05-in Wg Ð Air chg 15Ð17 changes/hr min ■ Central vacuum system 7. Safety ■ All electrical equipment is hazard-proof as defined in the National Electrical Code ■ Fire detection and suppression system (dry pipe, manual valve) 8. Security ■ Access control Ð KeyCard/cipher system Ð Intrusion detection system (BMS switches) Ð Vault doors with S&G 3-position tumbler Ð Lockable personnel and hardware access doors 9. Communications ■ Administrative phone ■ Operational Voice System (OVS) Cell 2 (Cells 1 and 3 planned) ■ Area warning system ■ Paging system t6

7-17 Table 7-4. Transfer Tower Area Table 7-5. Fairing Storage and Assembly Area Capability type Capability Capability type Capability ■ 1. Space/ ■ 27-ft by 30-ft clear floor access 1. Space/ Floor loading 75 psf on platforms ■ access ■ Design floor loading is 100 psf access 32-by 63-ft internal floor space ■ ■ Seven platforms on three sides (north, 68-ft 6-in-h by 22-ft-w breechload door east and south) opening Ð 75 psf loading on platforms 2. Handling ■ 75-ton stationary hoist ■ 2. Handling ■ 75-ton stationary hoist Hook height of 166 ft 6 in above floor ■ Hook height of 166 ft 6 in above floor elevation 69 in elevation 69 in ■ Speeds ■ Speeds Ð Hoist 0.5, 5.0 and 10 fpm Ð Hoist 0.5, 5.0 and 10 fpm ■ Pendant control at elevation 139 ft 0 in and ■ Pendant control at elevation 139 ft 0 in 165 ft 7 in and 165 ft 7 in. 3. Electrical ■ 110 VAC, utility power ■ 3. Electrical ■ Utility power Hazard-proof electrical equipment as Ð 110 VAC defined in the National Electrical Code ■ Hazard-proof electrical equipment as Articles 500Ð516 defined in the National Electrical Code ■ Multi-point grounding per MIL-STD-1542 Articles 500Ð516 4. Liquids ■ None ■ Static grounding reel 5. Pneumatics ■ Compressed air 125 psig 4. Liquids ■ None Ð 3/8-in. QD interface 6. Environment ■ 100,000 clean room capability 5. Pneumatics ■ Compressed air 125 psig Ð Inlet air Class 5000 Ð 3/8-in QD interface ÐTemp 70° ± 5° F 6. Environment ■ 100,000 clean room capability Ð RH 30Ð50% Ð Inlet air Class 5000 Ð Dif 0.05-in. Wg Ð Temp 70° ± 5° F Ð Air chg 10Ð12 changes/hr min Ð RH 30Ð50% ■ Central vacuum system Ð Dif 0.05-in. Wg 7. Safety ■ All electrical equipment is hazard-proof as Ð Air chg 10Ð12 changes/hr min defined in the National Electrical Code ■ Central vacuum system Articles 500Ð516 7. Safety ■ All electrical equipment is hazard-proof as ■ Fire detection and suppression system defined in the National Electrical Code 8. Security ■ Access control Articles 500Ð516 Ð KeyCard/cipher system ■ Fire detection and suppression system Ð Intrusion detection system (BMS 8. Security ■ Access control switches) Ð KeyCard/cipher system Ð Lockable personnel and hardware Ð Intrusion detection system (BMS access doors switches) 9. Communi- ■ Paging system Ð Vault doors with S&G 3-position cations tumbler t8 Ð Lockable personnel and hardware access doors 9. Communi- ■ Administrative phone 7.2.4.2 Technical Support Areas. Figures 7-17 cations ■ Operational Voice System (OVS) and 7-18 illustrate the plan views of the IPF, show- ■ Area warning system ■ Paging system ing Levels 89 and 101 of the Technical Support t7 the 75-ton cranes servicing the High Bay, the side. (Level numbers are defined in feet, with the Checkout Cells and the Transfer Tower Area; and SLC-6 launch mount defined as Level 100). These the Checkout Cell 5-ton cranes. Vehicles and equip- figures show room sizes as well as potential func- ment enter through the main entry door in the west tions. Note that the clean elevator can only be end of the Airlock. Personnel and support equip- accessed from the Technical Support side on Level ment access to the Checkout Cells is provided 89 through the airlock (for support equipment) or through the airlock on Level 89 of the Technical the clean change room. From the elevator, any level Support Area. There is also a personnel airlock on the Processing Side can be accessed. entry door on the south side of the Airlock. The level 69 Payload Processing Room (6902) is shown 7.3 SPACECRAFT TRANSPORT TO LAUNCH in Figure 7-16; there are also rooms available on SITE Levels 79, 89, 99, 109, 119 and 129. The rooms are After completion of preparations in one of the 4.9 by 7.0 m (16 by 23 ft). spacecraft Processing Facilities the flight-config- 7-18 Table 7-6. Payload Processing Room 6902 Table 7-7. Payload Processing Room (PPR) 8910 Capability type Capability Capability type Capability 1. Space/ ■ Processing/storage room 6902: 1. Space/ ■ Processing/Storage Room: access Ð 21 ft 5 in by 23 ft access Ð 35 ft 6 in by 30 ft 0 in and Ð 495 sq ft 23 ft 6 in by 23 ft 6 in ■ Door openings shall accommodate an 1495 sq ft. total envelope of 4 by 6 by 7 ft ■ Door openings shall accommodate an 2. Handling ■ None envelope of 4-ft by 6-ft by 7 ft 3. Electrical ■ 110 VAC, utility power 2. Handling ■ None ■ 120/208 VAC 3 ph 3. Electrical ■ 110VAC Utility power ■ Multi-point grounding per MIL-STD-1542 ■ 120/208VAC 3 ph ■ Hazard-proof electrical equipment as ■ Hazard proof electrical equipment as defined in the National Electrical Code defined in the National Electrical Code 4. Liquids ■ None 4. Liquids ■ None 5. Pneumatics ■ None 5. Pneumatics ■ None 6. Environment ■ 100,000 clean room capability 6. Environment ■ Temp 65Ð80 F Ð Inlet air Class 5000 RH N/A ÐTemp 70° ± 5° F Dif 0.5-in Wg Ð RH 30Ð50% Air Chgs 15 changes/hr min (goal) Ð Dif 0.05-in. Wg ■ Central Vacuum System Ð Air chg 15 changes/hr min 7. Safety ■ Fire detection and suppression system 7. Safety ■ Fire detection and suppression system 8. Security ■ Keycard/cipher access control 8. Security ■ Access control ■ Intrusion Detection system (BMS switches) Ð KeyCard/cipher system ■ Lockable personnel and hardware access Ð Intrusion detection system (BMS doors switches) 9. Communi- ■ None Ð Lockable personnel and hardware cations access doors t12 9. Communi- ■ None cations t9

H54527.2

5 TON CRANE ENVELOPE 75 TON CRANE ENVELOPE TRANSFER AIRLOCK HIGH BAY TOWER 30' x 100' 30' x 147' 5 TON CRANE AREA ENVELOPE 27' x 30'

35' x 44' 35' x 44' 35' x 44' CELL CELL CELL 1 2 3

SE STORAGE

PAYLOAD N PROCESSING CLEAN ELEVATOR ROOM 21' X 23' W E

Figure 7-16. California Spaceport—Processing Areas 7-19 H54525.1 CLEAN ELEVATOR

SECURITY OFFICE CLEAN ROOM AIR MEN'S LOCK BREAK ROOM SECURE ROOM OPS WOMEN'S PHE CLEAN ROOM FAX DRESSING AREA MGR ROOM ROOM 1495 SQ FT

SECURE OFFICE CLEAN ROOM CONFERENCE 1125 SQ FT MIC 988 SQ FT ROOM SECURE ROOM OFFICE 152 SQ FT

ELEVATOR N

W E 0 10 20 30 40 50 60 70 80 90 100

Figure 7-17. California Spaceport—Level 89 Technical Support Area H54526.1

CLEAN ELEVATOR

SECURE OFFICE MEN'S COMM 165 SQ FT SECURE OFFICE ROOM RANGE FIBRE OPTIC OFFICE 1352 SQ FT WOMEN'S TRANSMISSION SYSTEM SECURE 165 SQ FT ROOM (FOTS) INTERFACE OFFICE 171 SQ FT

OFFICE SECURE 182 SQ FT OFFICE OFFICE OFFICE 1156 SQ FT 1600 SQ FT 1621 SQ FT OFFICE 165 SQ FT

DIRTY ELEVATOR

N 0 10 20 30 40 50 60 70 80 90 100

W E

S Figure 7-18. California Spaceport—Level 101 Technical Support Area

ured spacecraft is installed in a transportation han- portation of the spacecraft to the launch pad. The dling can and moved to SLC-2 to be mated to the container (commonly called the handling can) can Delta II launch vehicle. MDA provides the trans- be configured for either three-stage or two-stage portation container (Figure 7-19) to support transmissions. The height of the handling can varies 7-20 3048 H54344.3 Shackle Access dia Platform 120 (Inside Skin)

Cover Load Capacity (17,800 lb)

Extension Ladder 1171 (Typ) 46.12 5639 222 3962 156

Handling Can (Shown with 5 Cylindrical Sections)

4115 Wheel Base 2946 162 116 Track Width

Conical Section 1294 for 3-Stage Missions 50.93

mm All dimensions are in Handling Can Configuration in. for 3-Stage Missions

3048 Shackle Access dia Platform 120 (Inside Skin)

Cover

1171 (Typ) 6915 PAF 46.12 (Ref)

AAAAA GSE Clamp AAAAA AA Payload (Ref) AAAAA AAAAA AAAAA AAAAA Handling Can (Shown with 4 Cylindrical AAAAA Sections) AAAAA

Direct Mate Adapter AAAAA for 2-Stage Missions Handling Can Configuration for 2-Stage Missions

Figure 7-19. Upper-stage Assembly Ground Handling Can and Transporter 7-21 according to the number of cylindrical sections vehicle. An aerial view of SLC-2 is shown in Figure required for a safe envelope around the spacecraft. 7-21. The spacecraft handling can is supported on a Because all operations in the launch complex rubber-tired transporter and, in a convoy, is slowly involve or are conducted in the vicinity of liquid or towed to the launch pad with MDA-provided trac- solid propellants and/or explosive ordnance de- tors and security personnel. The spacecraft con- vices, the number of personnel permitted in the area, safety clothing to be worn, type of activity tainer is purged with GN2 to reduce the relative humidity of the air inside the container and to permitted, and equipment allowed are strictly regu- maintain a slight positive pressure. Temperature in lated. Adherence to all safety regulations is re- the container is not controlled directly, but is main- quired. Briefings on all these subjects are given to those required to work in the launch complex area. tained at acceptable levels when transporting the The SLC-2 MST (Figure 7-22) is a 54.3-m (178- spacecraft by selecting the time of day at which ft) high structure with nine working levels desig- movement occurs. The Transportation Environ- nated as A, B, C, 1, 2, 3, 4, 5, and 6. An elevator ment is monitored with recording instrumentation. provides access to seven of the levels, B through 5. The white room (spacecraft area) 7.4 SPACE LAUNCH COMPLEX 2 encloses Levels 4, 5, and 6 (Figures 7-23 and 7 -24). SLC-2 (Figure 7-20) consists of one launch pad However, Level 4 is not used for spacecraft work. (SLC-2), a blockhouse, a Delta operations building, Levels 4 and 5 are fixed platforms, and Level 6 is an shops, a supply building, and other facilities neces- adjustable platform with a range of 399 cm (157 in.) sary to prepare, service, and launch the Delta (Figure 7-25). The white room enclosure is con-

H54345.2

1st & 2nd Stage Launch Pad Processing (HPF) MST Warehouse Support & (1615) Blockhouse Proofload FacilitiesAAA A(1622) A A A AAAAAAAAAAAAAAAAAAAAPump House N AAAAAAAAAAAAAAAAAAAAAA Engineering SupportAA (1628) AAAAAAAAAAAAAAAAAAAAAANASA - Hazardous A Processing Solid Rocket Motor Facility AAAAAAAAAAAAAAAAAAAAFacility (1610) CrossroadAA (1670) AAAAAAAAAAAAAAAAAAAATangair Astrotech Processing AAAAAAAAAAAAAAAAAAAAFacility (1032)

Figure 7-20. Space Launch Complex-2, Vandenberg Test Center 7-22 H54717.1 DAC1311768

Mobile Service Tower

RIFCA Alignment Bldg

Umbilical Mast

❏ Blockhouse ❏ Delta Operations Bldg

Storage Bldg

Figure 7-21. Space Launch Complex SLC-2, VAFB–Aerial View Looking East structed of RF-transparent panels. An internal bridge house, which is equipped with vehicle monitoring crane with a 4545-kg (5-ton) capacity, used for and control equipment. Space is also allocated for fairing and spacecraft equipment that must be moved use by spacecraft personnel (Figures 7-27 and 7- within the MST, has a maximum hook height of 28). In addition, a spacecraft console (Figure 7-29) 9.83 m (32 ft 3 in.) above Level 5 (Figure 7-26). that will accept a standard rack-mounted panel is Space is available on Level 5 for spacecraft GSE. available. Terminal board connections in the con- Placement of the GSE must be coordinated with sole provide electrical connection to the space craft MDA and appropriate seismic restraints provided. umbilical wires.

The entire MST is constructed to meet explosion- 7.5 SUPPORT SERVICES proof safety requirements. The restriction on the number of personnel admitted to the white room is 7.5.1 LAUNCH SUPPORT governed by safety requirements, as well as the For countdown operations, the launch team is limited amount of work space and the cleanliness located in the blockhouse at Space Launch Com- level required on the spacecraft levels. plex 2, and Buildings 836 and 840, with support Launch operations are controlled from the block- from other base organizations. 7-23 H54342.4

Elevation 177 ft, 11 in.

Lightning Rod

External Bridge Crane Mobile Service Tower Hook Height 152 ft, 0 in. Internal Bridge Crane Hook Height 143 ft, 3-1/2 in. Fixed Umbilical Tower FUT Lev 16 El 131 ft, 9 in. El 131 ft, 7-1/8 in. Level 6 Sta 320.12 FUT Lev 15 Level 6 Adjustable El 118 ft, 8 in. Level 6 Sta 477.12 FUT Lev 13 El 111 ft, 0-1/2 in. Level 5 El 110 ft, 7-1/8 in. Sta 568.62 El 103 ft, 3 in. FUT Lev 12 Sta 662.12 Level 4 FUT Lev 11 El 94 ft, 4 in. Level 3 El 94 ft, 7-1/8 in. Sta 769.12 FUT Lev 10 El 85 ft, 0 in. Sta 881.12 Level 2 FUT Lev 9 El 78 ft, 7-1/8 in.

MST Level 1 El 64 ft, 3-1/2 in. Sta 1129.62

MST Level C El 44 ft, 0 in. Sta 1373.12

MST Level B El 28 ft, 10 in. Sta 1555.12

MST Level A El 16 ft, 10 in. Sta 1699.12 Boattail El 14 ft, 11-1/8 in. Sta 1722 0 ft, 0 in. Ground Level

Figure 7-22. SLC-2 MST/FUT Elevation 7-24 H54340.3

All dimensions are in Meter Foot

Elevator

Front Sliding Doors Ladder Up to Level 6

Hatchway to Level 4 Hinged Line of Sight to Data Transfer Platforms Antenna at Building 836 10.5 (Typ) (148.5 deg From True North) deg

True North lV Fairing l Storage

5.5 8.7 7.5 deg 18.0 28.7

lll Downrange ll Down Up

AAAA 2.7/9.0 dia AAA

Grounded Aluminum Diamond Tread Plate 2.8 9.25 6.5 3.8 21.2 12.5

Notes: ❏ Downrange refers to the orientation of the ❏ launch pad and not the Delta trajectory 0 1.5 3 Scale in Meters ❏ The location of the spacecraft GSE on ❏ Level 5 must be coordinated with MDA

0246810 120 Volt, 60 Amp, Phase 1 Outlet Scale in Feet

Figure 7-23. Level 5 of SLC-2 Mobile Service Tower–Plan View 7-25 H54341.3

Meter All dimensions are in Foot

Fairing Storage

Travel Envelope of 5-Ton Capacity Front Sliding Doors Ladder From Level 5 Interior Bridge Crane

Landing at Elevation 123-ft 4-in.

Downrange

Line of Sight to Data Transfer Antenna at Building 836 (148.5 deg From True North) 10.5 deg Down Open Down True North IV to Level 5

Fairing Split I LIne

5.5 8.7 7.5 deg 18.0 28.7

Ill Il AAAAADownrangeAA 10-ft 10-in. With Inserts 12-ft 0-in. With Inserts Removed Self-Adjusting Stairway AAAAAAA 12 ftÐ0 in.

2.8 9.25 6.5 3.8 21.2 12.5 Grounded Aluminum Diamond Tread Plate

Notes: ❏ Downrange refers to the orientation of the ❏ launch pad and not the Delta trajectory 0 1.5 3 Scale in Meters 120 Volt, 60 Amp, Phase 1 Outlet

0246810 Scale in Feet

Figure 7-24. Level 6 of SLC-2 Mobile Service Tower–Plan View 7-26 02619REU7

Sta 203.99 mm All dimensions are in in. All station numbers are in inches Note: SLC-2 Level 6 can be adjusted within the range shown with the vehicle on stand

Level 6 (Max) Sta 320.12

10-ft Fairing

3988 Adjustable 157 Range of Level 6

3302 dia with Inserts Installed 130 3658 dia with Inserts Removed 144

Level 6 (Min)

Sta 477.12

2324 91.5

Sta 553.39 15.23-in. Level 5

Sta 568.62 2743 dia 108

Figure 7-25. Spacecraft Work Levels in SLC-2 Mobile Service Tower–VAFB (Download Figure) 7-27 H54338.5 Weather Enclosure

Enclosure Door Attached to Crane Bridge 20-ton Exterior Bridge Sliding Roof Hook Height 152 ft, 0 in.

5 ft, 10 in.

5-ton Sta 203.99 Interior Bridge 30 ft, 8 in. (max) Hook Height Elevation 143 ft, 3-1/2 in. (max) 10-ft dia

Level 6 (max) Elevation 131 ft, 9 in. Sta 320.12 Level 6 Landing Adjust 123 ft, 4 in. Third Stage/Spacecraft Container Level 6 (min) Elevation 118 ft, 8 in. Sliding Front Doors Sta 477.12 Level 5 Elevation 111 ft, 0-1/2 in. Sta 568.62

Figure 7-26. Whiteroom Elevations and Hook Heights—SLC-2 Mobile Service Tower

7.5.1.1 Mission Director Center (Building 840). 7.5.1.3 Launch Decision Process. The launch The Mission Director Center, Figure 7-7, provides decision process is made by the appropriate man- the necessary seating, data display, and communi- agement personnel representing the spacecraft, cations to control the launch process. Seating is launch vehicle, NASA, and range. Figure 7-30 provided for key personnel from MDA, Western shows the communications flow required to make Range, and the spacecraft control team. For NASA the launch decision. For NASA missions, a Mission launches, key NASA personnel will also occupy Director, launch management advisory team, engi- space in the Mission Director Center. neering team, and quality assurance personnel will 7.5.1.2 Space Launch Complex 2 Blockhouse. also participate in the launch decision process. Launch operations are controlled from the block- 7.5.2 Weather Constraints house, which is equipped with vehicle monitoring and control equipment. Space is also allocated for 7.5.2.1 Ground-Wind Constraints. The MST the spacecraft blockhouse consoles and console encloses the Delta II until approximately L-7 hours operators. Terminal board connections in the space- and provides protection to the vehicle from ground craft blockhouse junction box provide electrical winds. The winds are measured using an anemom- connection to the spacecraft umbilical wires. eter at the 16.5-m (54-ft) level of the MST. 7-28 H54612.2

m Paved Parking ft N

Access Ramp

A.C. W.C. Down

T/M Station

ALCS Control Room Lobby

Door Blocked Entrance

Observation Existing Spacecraft Comm Room A.C. Console Room Available for additional Spacecraft Consoles

1.5 0 3 DRIMS 2.4/8.0 m Control Room 1 4.3 0 5 10 14.0 ft

Figure 7-27. SLC-2 Blockhouse H30434.4

H30434.4 59 in.

Scale 0246 Talker/ (Feet) Timer Console Ramp Support I-Beam Up (Height Above Floor = 78 in.)

Launch Vehicle Technical Advisor Console Management (Removable)

50 ft

Launch Conductor, Safety, 1B16196-1 Customer Management Console Spacecraft Console Quality Assurance Booster Technical 12 ft 10 in. Monitor Control Second Stage Adviser 11 ft 3 in. Console Area for Locating Spacecraft Consoles AC Power Panel 20 ft Limits of Spacecraft Area Communications J-Box Cable Trench (Stay Out Area) Figure 7-28. SLC-2 Blockhouse Control Room 7-29 H54611.1

mm Spacecraft wiring is supplied by the Delta project to the in. spacecraft blockhouse console and terminated to a terminal strip. Users are required to supply the cable from their equipment in the console to the terminal strip — a distance of approximately 1219 mm (48 in.) — with lugs capable of accepting a 3.5-mm (0.138-in.) dia machine screw.

15.87 6.35 Panel Mounting 0.62 0.25 Hole Pattern (Typical Both Sides)

Communications Panel 333 52.5

15.87 0.62 12.70 0.50 546 15.87 21.5 0.62

584 610 23.0 24.0

400/15.75 AAAAATB2AATB2 Panel Space AAAAAAA AAAAAAA Console terminal block AAAAAAA (P/N AMP 601805-1) TB1/TB2 near side, TB3/TB4 far side. Standard AAAAAAA Spacecraft agency will 483/19.0 provide Burndy lugs Panel Width YAEV10-T7 (#12 AWG) and AAAAAAA YAE18N1 (#16 or #20 AWG) 457 18.0 AAAAAAA AAAAAAA AAAAAAA

Figure 7-29. Spacecraft Blockhouse Console-Western Range 7-30 H54604.3 Spacecraft Mission Director Spacecraft Ground Station Center (Bldg 840) Mission Control Center S/C Status Spacecraft S/C Network Spacecraft Spacecraft Spacecraft Spacecraft Network Status Status Network Project Manager Mission Director Status Manager (User) (User) Spacecraft (User) Ground Station Spacecraft (User) Vehicle Status Launch Launch Director of Vehicle Mission Concurrence Spacecraft Status Site Controller Launch Engineering Director Mission Advisory (NASA) Vehicle (MDA) (MDA) Control Center System (User) Status Launch Status Decision

Launch Vehicle Space Launch LOCC Ð Launch Systems Engineer Complex 2W Operations Control (MDA) Launch Center (Bldg 7000) Blockhouse Director (MDA) USAF Vehicle Status Status (30 SW/CC) Launch Vehicle TOPS 1 Data Center Range ROC (LVDC) Chief Field Launch Coordinator RCO (Bldg 836) Engineer Conductor (MDA) (MDA) (MDA) SMFCO (30 SW) ¥ Range Safety Status ¥ Western Range Spacecraft Status Coordinator Status (MDA) ¥ Weather ¥ Network Status

Figure 7-30. Launch Decision Flow for Commercial Missions–Western Range

The following ground-wind limitations includ- 7.5.2.2 Winds-Aloft Constraints. Measurements ing gusts apply: of winds aloft are taken in the vicinity launch pad. A. The MST shall not be moved from the Delta II The Delta II controls and loads constraints for winds if ground winds in any direction are predicted to aloft are evaluated on launch day by conducting a exceed 30 knots (35 mph) or if the actual measured trajectory analysis using the measured wind. A wind speed is in excess of 25 knots (30 mph) curvefit to the wind data provides load relief in the measured at the 16.5-m (54-ft) anemometer height. trajectory analyses. The curvefit and other load- B. The maximum allowable ground winds at the relief parameters are used to set the mission con- 31.1-m (102-ft) level are shown on Figure 7-31 for stants just prior to launch. 792X vehicles with lengthened nozzles on the air- 7.5.2.3 Weather Constraints. Weather constraints ignited GEMs. As noted on the figure, the con- are imposed to ensure safe passage of the Delta II straints are a function of the predicted liftoff solid launch vehicle through the atmosphere. The follow- motor propellant bulk temperature. This figure is ing is a general overview of the constraints evalu- applicable for both 9.5-ft and 10-ft diameter fairing ated prior to liftoff. Appendix B list the detailed configurations. This plot combines liftoff controls, weather constraints. liftoff loads, and on-stand structural ground-wind A. The launch will not take place if the normal restrictions. flight path will carry the vehicle: 7-31 H54518.3 Delta II 7925-10 Ground Wind Velocity Criteria Between 280¡ and 340¡ (92 ft) N Six GEM Solids off the Pad, Three GEM LN Solids Airlit WR-Launch Pads SLC-2 Temp (¡F) Wind Speed (Knots) 30 25 35 26 Vehicle Configuration 792X 40 27 Fairing Diameter 9.5- & 10-ft 45 28 Solids GEM LN 50 29 0 Launch Site WR 330 30 Minimum Solid Motor Propellant Bulk 30¡Ð50¡ F Temperature Range Anemometer Level 102 ft

300 60

No Launch Launch Angles Indicate Direction From Which Winds Come (Wind Speed is Measured at 102 Feet) W 27 50 40 30 20 10 0 10 20 30 40 50 90 E 0 Knots Pad Azimuth 259 deg

240 120

35 Knots (102 ft) Between 89¡ and 168¡ (92 ft) Temp (¡F) Wind Speed (Knots) 210 150 30 22 180 50 23 S

Figure 7-31. Delta II 792x Ground Wind Velocity Criteria, SLC-2

1. Within 18.5 km (10 nmi) of a cumulonimbus free line of sight exists to the vehicle flight path. (thunderstorm) cloud, whether convective or in Rawinsonde will not be used to determine cloud layers, where precipitation or virga is observed. buildup. 2. Through any cloud, whether convective or D. Even though the above criteria are observed, in layers, where precipitation or virga is present. or forecast to be satisfied at the predicted launch 3. Through any frontal or squall-line clouds time, the launch director may elect to delay the which extend above 3048 m (10,000 ft). launch based on the instability of the current atmo- 4. Through cloud layers or through cumulus spheric conditions. clouds where the freeze level is in the clouds. 5. Through previously electrified clouds not 7.5.2.4 Lightning Activities. The following are monitored by an electrical field mill network if the procedures for test status during lightning activity: dissipating state was short-lived (less than 15 A. Evacuation of the MST and the fixed umbilical minutes after observed electrical activity). tower (FUT) is accomplished at the direction of the B. The launch will not take place if there is Launch Conductor (Reference: Delta Launch Com- precipitation over the launch site or along the flight plex Safety Plan). path. B. First- and second-stage instrumentation may C. A weather observation aircraft is mandatory be operated during an electrical storm. to augment meteorological capabilities for real- C. If other vehicle electrical systems are powered time evaluation of local conditions unless a cloud- when an electrical storm approaches, these systems 7-32 may also remain powered. If category A electro- Authority List (EAL), and they must sign in and explosive device (EED) circuits are electrically exchange badges at the access control station. connected in the launch configurations, the guid- 7.5.4.2 Hazardous Processing Facility. Physi- ance computer must be turned off. cal security at the Hazardous Processing Facility D. If an electrical storm passes through after a (Building 1605, 1610) is provided by a chain link simulated flight test, all electrical systems are turned perimeter fence, a guard house gate, door locks, and on in a quiescent state, and all data sources are guards. Details of the spacecraft security require- evaluated for evidence of damage. This turn-on is ments are arranged through the MDA spacecraft done remotely (pad clear) if any category A ord- coordinator. nance circuits are connected for flight. Ordnance 7.5.4.3 Spacecraft Processing Laboratories. circuits are disconnected and safed prior to turn-on Physical security at the Spacecraft Processing Labo- with personnel exposed to the vehicle. ratories (Building 836) is provided by door locks E. If data from the quiescent turn-on reveal and guards. Details of the spacecraft security re- equipment discrepancies that can be attributed to quirements are arranged through the MDA space- the electrical storm, a flight program requalification craft coordinator or appropriate payload processing test must be run subsequent to the storm and prior to facility. a launch attempt. 7.5.4.4 VAFB Security. For access to VAFB, US F. During terminal countdown, the launch direc- citizens must provide full name with middle initial tor is responsible for initiating and ending an alert. if applicable, social security number, company name, Upon initiation of an alert, the GC is turned off. and dates of arrival and expected departure to the When the alert is lifted, the GC is turned on and MDA spacecraft coordinator or MDA/VAFB memory verified. Security. MDA Security will arrange for entry

7.5.3 Operational Safety authority for commercial missions or individuals sponsored by MDA. Access by NASA personnel or Safety requirements are covered in Section 9 of US Government sponsored Foreign Nationals is this manual. In addition, it is the operating policy at coordinated by NASA KSC with the USAF at MDA that all personnel will be given safety orien- VAFB. For non-United States citizens, clearance tation briefings prior to entrance to hazardous areas information (name, nationality/citizenship, date and such as SLC-2. These briefings will be scheduled place of birth, passport number and date/place of by the MDA spacecraft coordinator and presented issue, visa number and date of expiration, and title by the appropriate safety personnel. or job description) must be furnished to MDA two weeks prior to the VAFB entry date. Government 7.5.4 Security sponsored individuals must follow NASA or U.S. 7.5.4.1 Space Launch Complex 2. SLC-2 secu- Government guidelines as appropriate. The space- rity is ensured by perimeter fencing, interior fenc- craft coordinator will furnish visitor identification ing, guards, and access badges. White room access documentation to the appropriate agencies. After is controlled at the MST entrance. To gain access, MDA Security gets clearance approval, entry to visitors’ names must be on the White Room Entry VAFB will be the same as for US citizens. 7-33 7.5.5 Field-Related Services The countdown schedules provide a detailed MDA employs certified propellant handlers en- hour-by-hour breakdown of launch pad operations, semble (PHE) suit propellant handlers, equipment illustrating the flow of activities from spacecraft drivers, welders, riggers, and explosive ordnance erection through terminal countdown, and reflect- handlers, in addition to people experienced in most ing inputs from the spacecraft project. These sched- electrical and mechanical assembly skills, such as ules comprise the integrating document to ensure torquing, soldering, crimping, precision cleaning, timely launch pad operations. and contamination control. MDA has under its Typical schedules of integrated activities from control a machine shop, metrology laboratory, LO2 spacecraft weighing in the HPF until launch (Fig- cleaning facility, and proof-loading facility. MDA’s ures 7-33 through 7-45) are shown as launch minus operational team members are familiar with USAF (F-) workdays. Saturdays, Sundays, and holidays and NASA payload processing facilities at VAFB are not scheduled workdays and, therefore, are not and can offer all of these skills and services to the F- days. The F- days, from spacecraft mate through spacecraft project during the launch program. launch, are coordinated with each spacecraft agency to optimize on-pad testing. All operations are for- 7.6 DELTA II PLANS AND SCHEDULES mally conducted and controlled using launch count-

7.6.1 MISSION PLAN down documents. The schedule of spacecraft activities during that time is controlled by the MDA A Mission Plan (Figure 7-32) is developed for Launch Operations Manager. Tasks involving the each launch campaign showing major tasks on a spacecraft or tasks requiring that spacecraft person- weekly timeline format. The plan includes launch nel be present are shaded for easy identification. vehicle activities, prelaunch reviews, and space- A typical three-stage mission from VAFB is as craft processing area occupancy times. follows; spacecraft and third-stage checkout are 7.6.2 Integrated Schedules completed before F-12 day. The schedule of spacecraft activities before inte- F-12 Tasks include equipment verification, pre- grated activities in the HPF varies from mission to cision weighing of spacecraft, and securing. mission. The extent of spacecraft field testing var- F-11 Spacecraft is lifted and mated to the third ies and is determined by the spacecraft agency. stage. The clampband is installed and the initial Spacecraft/launch vehicle schedules are similar clampband tension established. from mission to mission from the time of spacecraft F-10 Final preparations are made prior to can-up weighing until launch. for both spacecraft and third stage, and spacecraft/ Daily schedules are prepared on hourly timelines third stage interface verification is done, if required. for these integrated activities. These schedules cover F-9 The payload handling can is assembled four days of integrated effort in the HPF and 8 days around the spacecraft/third stage, and handling-can of launch countdown activities. HPF tasks include transportation covers are installed. The can is placed spacecraft weighing, spacecraft third stage mate on its trailer and its purge is set up. and interface verification, and transportation can F-8 Tasks include transportation to the launch assembly around the combined payload. site, erection and mating of the spacecraft/upper

7-34 H30438.2 JANUARY FEBRUARY MARCH APRIL MAY 2027 4 11 18 25 1 8 7 14 21 28 7 14 21 28 4 11 18 25 2 9 16 23 30 DMCO C/O 21 SHIP UPPER STG BUILDUP PRESHIPAAAAA MEETINGAAAAA - AT PUEBLOAAAHPTF AAAAAA (BLDG 1610) FAIRING PROCESSING( BLDG 1610) SOLID MOTOR BUILDUP (BLDG 1670) AAAAAAAAAAAAAINTERSTAGE PROCESSINGAA FIRST STAGE AT HPF LEGEND LV OPS SECOND STAGEAA AT HPFAA AA VOS READINESS REVIEW 29 S/C OPS FIRST STAGE ERECTION INTERSTAGE ERECTIONAAA INTERSTAGE MATE FAIRING ERECTION SECOND STAGE ERECTION AND MATEAAA GEMS S/M ERECTION PRETEST PREPS VEH AASYSTEMSAAA CHECKOUTAAAA S/C PROCESS SPACECRAFT AVAILABLE 25 SPACECRAFT PREMATE REVIEW 25 LAUNCH SITE READINESS REVIEW 4 PAYLOAD ERECTION 5 S/C SYSTEMS TEST (5/6) FLT PROG VERIF (5/9)A ORDNANCE INSTL (5/10) FAIRING INSTL (5/11-5/13)AAA FLIGHT READINESS REVIEW 13 SS PROP LOAD TEST (5/16) GC, RS, BEACON CHECKS (5/17)A INTEGRATED OPERATIONS LAUNCH READINESSAAA REVIEWA 18 LAUNCH 19 Figure 7-32. Typical Mission Plan

H30439.5 0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300

WEIGH SPACECRAFT BRIEFING AT BLDG 1610

BAY OPENING CHECKS

SETUP/CHECKOUT PWU

HOIST FUNCTIONAL/STRAY VOLTAGE CHECKS

LEGEND POSITION CLASS C WEIGHTS PAD OPEN WEIGH S/C ITEMS TO BE INSTALLED LATER

A FLASHING AMBERÐ HYDRASET/LOAD CELL LINKAGE SETUP A LIMITED ACCESS FLASHING REDÐ LOAD CELL SHUNT CHECKS PAD CLOSED A CLASS C WEIGHT LIFT (VERIFY REPEATABILITY) S/C ACTIVITY SPACECRAFT LIFT/WEIGH/LOWER *

SPACECRAFT LIFT/WEIGH/LOWER *

SECURE LIFT EQUIPMENT * NOTE: LIFT AND LOWERING STEPS TO BE * ACCOMPLISHED BY S/C PERSONNEL SECURE WEIGH EQUIPMENT

BALLAST WEIGHTS (IF REQD)

Figure 7-33. Typical Spacecraft Weighing (F-12 Day) 7-35 H30440.6 0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300 VISHAY EQUIPMENT WARMUP S/C PAM MATE BRIEFING AT BLDG 1610

BAY OPENING CHECKS

LEGEND VISHAY INSTRUMENT STUD CALIBRATIONS AAPAD OPEN ACTUATOR INSTALLATION & LOCKWIRE FLASHING AMBERÐ CLAMPBAND PREPARATIONS AALIMITED ACCESS FLASHING REDÐ HOIST STRAY VOLTAGE & CRANE FUNCTIONAL TESTS PAD CLOSED AA LIFT/TRAVERSE/MATE SPACECRAFT S/C ACTIVITY SPACECRAFT TO PAF GAP MEASUREMENTS

CLAMPBAND INSTALLATION

COMPLETED PRIOR TO THIS DATE: CLAMPBAND TENSIONING/TAPPING ¡ CLAMP BAND DETAIL INSPECTION/LUBRICATION ¡ ENGINEERING WALKDOWNS SECURING ¡ PHOTOGRAPH DOCUMENTATION ¡ WORK STAND CLEAN/MOVE INTO POSITION VISHAY RECHECKS ¡ PAF/SPACECRAFT INTERFACE INSPECTION S/C PAM INTERFACE VERIFICATION (IF REQD)

Figure 7-34. Typical Spacecraft/PAM Mate (F-11 Day)

H30441.6 0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300

S/C PAM FINAL PREPARATIONS BRIEFING AT BLDG 1610 BAY OPENING CHECKS SEPARATION CLAMP BAND FINALIZING

GAP MEASUREMENTS END FITTINGS

LEGEND INSTALL BAND RETAINERS PAD OPEN CONNECT SPRINGS TO RETAINERS

FLASHING AMBERÐ CONNECT/TORQUE ETA INTO CUTTERS AALIMITED ACCESS INSTALL ATTACH BOLT CUTTER BRACKETS AAFLASHING REDÐ PAD CLOSED LOCKWIRE SHIELDS/BRACKETS ETA

S/C ACTIVITY INSTALL NON-FLIGHT TAGS

SEPARATION BLANKET INSTL.

FINAL INSPECTION

PHOTOGRAPH ASSEMBLY

TRAILER PURGE SETUP

CLEAN & PREASSEMBLE CYLINDRICAL SECTIONS OF TRANSPORT CAN

INSTALL/TORQUE 4 TRANSPORT CAN RING ASSEMBLIES TO SPIN TABLE

Figure 7-35. Typical Spacecraft/PAM Final Preparations (F-10 Day) 7-36 H30442.4 0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300 TRANSPORTATION CAN INSTALLATION BRIEFING AT BLDG 1610 BAY OPENING CHECKS

ENGINEERING WALKDOWN LEGEND CRANE FUNCTIONAL CHECKS PAD OPEN A CRANE STRAY VOLTAGE CHECKS FLASHING AMBERÐ LIMITED ACCESS A AHOIST INSPECTION FLASHING REDÐ EQUIPMENT PROOF LOAD VERIFICATION A PAD CLOSED

S/C ACTIVITY INSTALL CONICAL SEALS INSTALL HUMIDITY/TEMP RECORDER (IF REQD) COMPLETED PRIOR TO THUS DATE: INSTALL CYLINDER SHELLS ¡ CLEAN/ASSEMBLE/TRANSPORT CAN ¡ CLEAN/DISASSEMBLE/PREP CAN BAG THE CAN ASSEMBLY ¡ FAB ILLUMALLOY BAG ¡ CLEAN/MOVE TRAILER IN BAY ¡ 4 RING ASSY MATED TO SPINTABLE REMOVE NOZZLE THROAT PLUG

LIFT S/C & PAM/MATE TO TRAILER

TRAILER PURGE SETUP

PURGE CAN ASSEMBLY

ATTACH IMPACT RECORDER

Figure 7-36. Typical Transportation Can Installation (F-9 Day)

H30443.5 0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300 TRANSPORTATION BRIEFING AT BLDG 1610

TRANSPORT SPACECRAFT FROM BLDG 1610

ERECTION PREPS

ERECTION BRIEFING UNDER THE HOOK AT SLC-2W

LEGEND ERECT & MATE SPACECRAFT TO SS PAD OPEN UNCAN SPACECRAFT AFLASHING AMBERÐ LIMITED ACCESS DEERECT S/C CAN AFLASHING REDÐ APAD CLOSED WHITEROOM STABILIZATION S/C ACTIVITY BOLT PAF TO SECOND STAGE

S/C TO INSTALL CABLE UP PAM STAGE B/H CONSOLE 3 DAYS PRIOR TO INITIATE S/C BATTERY CHARGING THIS DAY EMERGENCY POWER TEST

VEHICLE COUNTDOWN PREPS

BATTERY PREPS & INSTALLATION

Figure 7-37. Typical Spacecraft Erection (F-8 Day) 7-37 H30444.6 0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300

DATA STATION WARMUP GUIDANCE SECTION AIR ON PRETEST BRIEFING FOR FLIGHT PROGRAM VERIF TEST POWER ON & PRETEST PREPS S/C POWER ON & DRESS REHEARSAL AZIMUTH DETERMINATION PREPS LEGEND COMM CHECK PAD OPEN A MINUS COUNT (ABBREV.TERM.COUNT) FLASHING AMBERÐ SPACECRAFT POWER IN LAUNCH MODE(T-4 MIN.) LIMITED ACCESS A T-0 FLASHING REDÐ S/C PRELAUNCH MODE APAD CLOSED + COUNT (FLT PRG VERIF TEST) S/C ACTIVITY TEST RECYCLE & CIRCUIT BREAKER VERIFY POWER ON STRAY VOLTAGE AZIMUTH DETERM. & MONUMENT CKS TEST RECYCLE & BATTERY CONNECT ENGR W/D, PHOTOS & PART GUID SECT C/LO ENGR W/D, PHOTO'S & PART CENT SECT CL/O S&A C/O S/C BATTERY CHARGING AAA

Figure 7-38. Typical Flight Program Verification and Safety Stray Voltage Checks (F-7 Day)

H30445.3 0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300 PRETEST BRIEFING

PREPS

AARECEIVE S&A'S

AAAAAS&A INSTL & ROTATION CHECK (FS AND SS)

AAAAAAA PWR OFF STRAY VOLT& ORD CONNECT

AAAAFS S/M ENG W/D

FS BOATTAIL C/O &PREP FOR ETA HOOKUP

LEGEND CENTER SECT ENG W/D AAPAD OPEN FLASHING AMBERÐ CENTER SECT CLOSEOUT AALIMITED ACCESS FLASHING REDÐ MINISKIRT ENG WALKDOWN AAPAD CLOSED S/C ACTIVITY INSTALL SUPPORT PEDESTALS ON LEVEL 5 S/C BATTERY CHARGING

Figure 7-39. Typical Ordnance Installation (F-6 Day) 7-38 H30446.6 0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300 SS & PAF INSPECTION, STAGE CLEANING, REMOVE CATCH NETS & RIFCA INSULATION CLOSEOUT PHOTOS REMOVE S/C BAG & PURGE (MDA CRANE OPERATOR) LEGEND HOIST FUNCTIONALS & HOIST BEAM / FAIRING CONNECTION PAD OPEN S/C PREPS FOR FAIRING INSTL AFLASHING AMBERÐ FINAL S/C ACCESS PRIOR TO FAIRING INSTL LIMITED ACCESS FAIRING CLEANING (IF REQUIRED) AFLASHING REDÐ S/C ENG WALKDOWN PAD CLOSED PAF ENG WALKDOWN A S/C ACTIVITY RAISE LVL 6 FAIRING INSTL BRIEFING INSTALL FAIRING SECTOR III IN PEDESTAL (RESTRAIN AT LVL 6) INSTL FAIRING SECTOR II IN PEDESTAL, RESTRAIN AT LVL 6, INSTL GUIDE DEVICES, INSTL FAIRING SECTOR I IN PEDESTAL, LOWER LVL 6 INSTALL RESTRAINT AND GUIDE DEVICES ON LVL 6 INSTALL HALO AND MATE FAIRING SECTORS INSTALL FIELD JOINT BOLTS CLOCK FAIRING (IF REQUIRED) ESTABLISH FAIRING AIR COND (TEMPORARY) RAISE LVL 6 REMOVE STRONGBACKS SPACECRAFT BATTERY TRICKLE CHARGE Figure 7-40. Typical Fairing Installation (F-5 Day) H30447.4

0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300

FAIRING ELECTRICAL CONNECTIONS

DISCONNECT FAIRING A/C

RAISE LVL 6

REMOVE STRONGBACKS

LOWER LVL 6

INSTALL FAIRING A/C LEGEND BRIEFING PAD OPEN S.S. SERVICING PREPS AND BAS PREPS A FLASHING AMBERÐ LIMITED ACCESS SETUP A-50/N204 SENSOR SYSTEM

A FLASHING REDÐ PROP CHILL A PAD CLOSED S/C ACTIVITY S/C LIMITED FUNCTIONAL CHECK

FLIGHT READINESSS REVIEW

AIR CONDITIONING WATCH

SPACECRAFT BATTERY TRICKLE CHARGE

Figure 7-41. Typical Fairing Finaling (F-4 Day) 7-39 H30448.4 0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300

FAIRING CLEAT INSTL

FAIRING WEDGE INSTL LEGEND PAD OPEN

FLASHING AMBERÐ BAGGIE INSP & ELECT CHECKS (OFF PAD) A LIMITED ACCESS FLASHING REDÐ A PAD CLOSED PROP CHILL

S/C ACTIVITY

REDLINE OBSERVER BRIEFING

FAIRING ACOUSTIC SHIELD INSTL

SPACECRAFT BATTERY TRICKLE CHARGE

AIR CONDITIONING WATCH

Figure 7-42. Typical Fairing Finaling (F-3 Day) H30449.3 0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300

AIR CONDITIONING WATCH (V52T1) & S/C PURGE

PROP CHILL

DATA STATION WARMUP AAA FAIRING ORDN BRIEFING AAAINSTL AABAGGIE INSP & ELECT CHECKS (OFF PAD) AAFINAL PROP SERV PREPS & FINAL BAS PREPS LEGEND OXIDIZER LOAD PAD OPEN AAA LUNCH BREAK FLASHING AMBERÐ A LIMITED ACCESS AAA FUEL LOAD A FLASHING REDÐ AAAAA PAD CLOSED 2ND STG PROP SECURE

S/C ACTIVITY PROP AAWATCH

S/C MISSION REHEARSAL

S/C BATTERY CHARGING

Figure 7-43. Typical Second-Stage Propellant Loading (F-2 Day) 7-40 H30450.4 0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300 DATA STATION WARMUP LAUNCH READINESS REVIEW BRIEFING PREPS HEAT RP-1 S/C FUNCTIONAL CHECKS LEGEND S/C CLOSEOUT ACTIVITY PAD OPEN AZIMUTH PREPS 1ST & 2ND STAGE TURNON AAFLASHING AMBERÐ LIMITED ACCESS COMM CHECK

AAFLASHING REDÐ SLEW CHECKS PAD CLOSED BEACON CHECKS AA AZIMUTH UPDATE S/C ACTIVITY SECURING 2ND STAGE ENGINEERING WALKDOWN PROP SYS PREPS & LINE LVL RP-1 2ND STAGE THERMAL BLNKT INSTL ENGINEERING WALKDOWN I/S CLOSEOUT RANGE SAFETY CHECKS

CLASS A ORDN HOOKUP AAA RED TAG INVENTORY PREP FOR TOWER REMOVAL

A/C WATCH, PROP WATCH & SPACECRAFT BATTERY CHARGING

Figure 7-44. Typical Beacon, Range Safety, and Class A Ordnance (F-1 Day) H30451.4 0100 0300 0500 0700 0900 1100 1300 1500 1700 1900 2100 2300

AIR CONDITIONING & PROP WATCH HEATED RP-1 RECIRCULATE S/C FINAL CLOSEOUTS DESTRUCT CHECKS FINAL S/C ACCESS PRIOR TO LAUNCH BRIEFING CAMERA SETUP (PHOTO SQUADRON) LEGEND ENGINEERING WALKDOWN PAD OPEN PROP & PNEU FINAL CHECKS AIR COND SETUPS AFLASHING AMBERÐ LIMITED ACCESS FAIRING AND WHITEROOM PREPS LANYARD TENSIONING AFLASHING REDÐ PAD CLOSED MST REMOVAL & SECURING S/M ARMING BRIEFING AS/C ACTIVITY S/M ARMING AA PAD SECURING S/C COUNTDOWN

PRESS SS HE & N2 TO 1100 PSI&HEX FILL LCE PREPS & HOLD FIRE CHECKS BUILT-IN HOLD (60 MIN) S/C BATTERY TRICKLE CHARGING AAAAATERM. COUNT

Figure 7-45. Typical Delta Countdown (F-0 Day) 7-41 stage to the Delta II vehicle in the MST cleanroom, 7.6.3 Spacecraft Schedules and removal of the handling can from the tower. The spacecraft project will supply schedules to F-7 The flight program verification test is per- the MDA spacecraft coordinator, who will arrange formed, followed by the vehicle power-on stray support as required. voltage test. Spacecraft systems to be powered at liftoff are turned on during the flight program veri- 7.7 DELTA II MEETINGS AND REVIEWS fication test and all data are monitored for electro- During the launch scheduling preparation, vari- magnetic interference (EMI) and radio frequency ous meetings and reviews take place. Some of these interference (RFI). All spacecraft systems that will will require user input while others allow the user to be turned on at any time between F-7 day (stray monitor the progress of the overall mission. The voltage checks) and F-0 day (spacecraft separation) MDA spacecraft coordinator will ensure adequate will be turned on in support of the vehicle power-on user participation. stray voltage test. Spacecraft support of these vehicle system tests is critical in meeting the sched- 7.7.1 Meetings uled launch date. They have priority over other Delta Status Meetings. Status meetings are gen- spacecraft testing. erally held twice a week. They include a review of F-6 Tasks include Delta II vehicle ordnance the activities scheduled and accomplished since the installation/connection and preparation for fairing last meeting, a discussion of problems and their installation. solutions, and a review of the mission schedule. F-5, 4, 3 Spacecraft final preparations are Spacecraft representatives are encouraged to attend made prior to fairing installation, including Delta II these meetings. upper-stage closeout, preparations for second-stage Daily Schedule Meetings. Daily schedule meet- propellant servicing, and fairing installation. ings are held to provide the team members with F-2 Propellant is loaded into the second stage, their assignments and to summarize the previous or and fairing ordnance is installed. current day’s accomplishments. These meetings F-1 Tasks include launch vehicle guidance turn- are attended by the launch conductor, technicians, on, C-band beacon readout, guidance system azi- inspectors, engineers, supervisors, and the space- muth update, range safety checks, class A ordnance arming, final fairing preparations for MST removal, craft coordinator. Depending upon testing activi- second-stage engine closeout, launch vehicle final ties, these meetings are held at either the beginning preparations, and preparations for tower removal. or the end of the first shift. F-0 Launch day preparations include final space- 7.7.2 Prelaunch Review Process craft closeouts and fairing door installation, gantry Periodic reviews are held to ensure that the space- removal, final arming, terminal sequences, and launch. Spacecraft should be in launch configura- craft and launch vehicle are ready for launch. The tion immediately prior to T-4 minutes and standing Mission Plan (Figure 7-32) shows the relationship by for liftoff. The nominal hold and recycle point is of the review to the program assembly and test flow. T-4 minutes. Launch is typically scheduled for a The following paragraphs discuss the Delta II Thursday. readiness reviews. 7-42 Postproduction Review. This meeting, con- the activities since the pre-VOS review and verifies ducted at Pueblo, Colorado, reviews the flight hard- the readiness of the launch vehicle, third stage, ware at the end of production and prior to shipment launch facilities, and spacecraft for transfer of the to VAFB. spacecraft to the pad. Mission Analysis Review. This review is held at Flight Readiness Review (FRR). This review is Huntington Beach, California, approximately three an update of actuals since the Pre-VOS and is months prior to launch to review mission specific conducted to determine that checkout has shown drawings, studies, and analyses. that the launch vehicle and spacecraft are ready for Pre-Vehicle-On-Stand (VOS) Review. This countdown and launch. Upon completion of this review is held at Huntington Beach subsequent to meeting, authorization to proceed with the loading the completion of Delta mission checkout (DMCO), of second-stage propellants is given. This review and prior to erection of the vehicle on the launch also assesses the readiness of the range to support pad. It includes an update of the activities since launch and provides a predicted weather status. Pueblo, the results of the DMCO processing, and Launch Readiness Review (LRR). This review any hardware history changes. is held on L-1 day and all agencies and contractors Launch Site Readiness Review (LSRR). This are required to provide a ready-to-launch statement. review is held prior to erection and mate of the Upon completion of this meeting, an okay to enter upper stage and spacecraft. It includes an update of terminal countdown is given.

7-43 Section 8 is located at the MDA facility in Huntington Beach, SPACECRAFT INTEGRATION California. The prime objective of mission integra- This section describes the payload integration tion is to coordinate all interface activities required process, the supporting documentation required to launch commercial and government spacecraft via the Delta II launch vehicle. This objective from the payload agency, and resulting analyses includes reaching a customer-MDA interface agree- provided by McDonnell Douglas Aerospace. ment and accomplishing interface planning, coor- 8.1 INTEGRATION PROCESS dinating, scheduling, control, and targeting. The integration process developed by MDA is The Delta Program Office assigns a mission designed to support the requirements of both the integration manager to carry out interface activi- launch vehicle and the payload. We work closely ties. The mission integration manager develops a with our customers to tailor the integration flow to tailored integration planning schedule for the Delta meet their individual requirements. The integration II/spacecraft by defining the documentation and process (Figure 8-1) encompasses the entire life of analysis required. The mission integration man- the launch vehicle/spacecraft integration activities. ager also synthesizes the spacecraft requirements At its core is a streamlined series of documents, and engineering design and analysis into a con- reports, and meetings that are flexible and adapt- trolled mission specification that establishes able to the specific requirements of each program. agreed-to interfaces. Mission integration for commercial missions is The integration manager ensures that all lines of the responsibility of the Delta Program Office which communication function effectively. To this end,

H54819.1 Authotity to Proceed Requirements Definition

Release Initial Mission Specification

Production Planning ❏ Review/Study Spacecraft Requirements ❏ Engineering Compatibility Analysis ❏ Loads/Thermal/Mission/Controls ❏ Joint Agreements

Environment Test Plans

Fabrication Range Safety Documentation Range Network Documentation Assembly and Checkout Flight Software Launch Processing

Flight Readiness Reviews Launch Operations

Figure 8-1. Mission Integration Process 8-1 all pertinent communications, including technical/ 8.2 DOCUMENTATION administrative documentation, technical interchange Effective integration of the spacecraft into the meetings (TIM), and formal integration meetings, Delta launch system requires the diligent and timely are coordinated through the Delta Program Office preparation and submittal of required documenta- and executed in a timely manner. These data ex- tion. When submitted, these documents represent change lines not only exist between the user and the primary communication of requirements, safety MDA, but also include other agencies involved in data, system descriptions, etc., to each of the sev- Delta II launches. For NASA payloads, mission eral support agencies. The Delta Program Office integration activities are co-chaired by NASA acts as the administrative interface for proper docu- Goddard Space Flight Center (GSFC) Orbital Launch mentation and flow. All data, formal and informal, Services (OLS) and MDA. NASA OLS will assign are routed through this office. Relationships of the a mission integration manager to act as the point of various categories of documentation are shown in contact for the government. Figure 8-2 shows the Figure 8-3. typical relationships among agencies involved in a A general model for Delta Launch Vehicle and Delta mission. spacecraft documentation requirements is shown in

H30572.7

Spacecraft Spacecraft Orbital Network User Agency Support

NASA/GSFC*

Launch Vehicle Spacecraft Processing MDA Processing Facilities Facilities and Services Delta Program Office and Services

NASA USAF FAA GSFC KSC* ER/WR SMC

Data Network Launch Facilities Launch Facilities Delta II Procurement Licensing and Base Support and Base Support Support Quality Assurance Safety Certification (as Required) Spacecraft Processing Quality Assurance Facilities and Services Safety Surveillance MDA Communication and Data Support Range Safety and *For NASA Missions Only Ascent Tracking Quality Assurance

Safety Surveillance Data Network Support (as Required)

Figure 8-2. Typical Delta II Agency Interfaces 8-2 H30571.2

Spacecraft Requirements ❏ Spacecraft Questionnaire

Safety Compliance Integration Planning Mission Support ❏ Missile Systems Prelaunch Safety ❏ Operations❏ Schedule ❏ Package (MSPSP) ❏ Documentation❏ Reviews ❏ Operations Requirement (OR)/ ❏ Program Requirements Document (PRD) ❏ ¥ Range and Network Support ❏ Mission Support Request (MSR) ❏ Launch Operations Plan (LOP) Mission Specification ❏ Spacecraft and Vehicle Description ❏ Performance Requirements ❏ Interface Definition Launch Support ❏ ¥ Spacecraft/Delta ❏ Launch Processing Requirements ❏ ¥ Spacecraft/Fairing ❏ ❏ Payload Processing Requirements Vehicle/GSE (Mission Peculiar) ❏ Document (PPRD) ❏ Mission Compatibility Drawing ❏ ❏ Launch Site Test Plan (LSTP) Spacecraft to Blockhouse Wiring ❏ Integrated Procedures Diagram ❏ Launch Processing Documents (LPD)

Environmental Test Plans Mission Analysis ❏ Preliminary Mission Analysis (PMA) ❏ Spacecraft Qualification Verification ❏ ¥ Event Sequencing ❏ ¥ Ground Monitor and Tracking Overlay ❏ Detailed Test Objectives (DTO)

Figure 8-3. Typical Document Interfaces

Tables 8-1 and 8-2, and Table 8-3 describes the associated L-date (weeks before launch). The re- documents identified. Specific schedules are estab- sponsible party for each data item is identified. lished by agreement with each customer. The Close coordination with the Delta mission integra- Spacecraft Questionnaire shown in Table 8-4 is to tion manager is required to provide proper planning be completed by the spacecraft agency at least 2 of the integration documentation. years prior to launch to provide an initial definition of spacecraft characteristics. Table 8-5 is an outline 8.3 LAUNCH OPERATIONS PLANNING of a typical spacecraft launch-site test plan that The development of launch operations, range describes the launch site activities and operations support, and other support requirements is an evo- expected in support of the mission. Orbit data at lutionary process that requires timely inputs and burnout of the final stage is needed to reconstruct continued support from the spacecraft agency. The the performance of the Delta following the mission. relationship and submittal schedules of key con- A complete set of orbital elements and associated trolling documents are shown in Figure 8-5. estimates of 3σ accuracy required to reconstruct this performance is presented in Table 8-6, Data Re- 8.4 SPACECRAFT PROCESSING quired for Orbit Parameter Statement. REQUIREMENTS A typical integration planning schedule is shown The checklist shown in Table 8-7 is provided to in Figure 8-4. Each data item in Figure 8-4 has an assist the user in identifying the requirements at 8-3 Table 8-1. Spacecraft Agency Data Requirements each processing facility. The requirements identi- Nominal Due weeks fied are submitted to MDA for the PRD. MDA Table 8-3 Ð or + Description reference launch coordinates with CCAS/KSC, Astrotech, Califor- Spacecraft Questionnaire 2 L-104 nia Spaceport, or VAFB/KSC as appropriate and FAA License Information 2 L-104 Spacecraft Mathematical Model 3 L-90 implements the requirements through the PRD/ Spacecraft Environmental Test 5 L-84 Documents PPRD. The user may add items to the list. Please Mission Specification 4 30 days note that most requirements for assembly and check- Comments after receipt Electrical Wiring Requirements 7 L-80 out of commercial spacecraft will be met at the Spacecraft Drawings (Init/Final) 18 L-78/L-44 Fairing Requirements 8 L-68 Astrotech or California Spaceport facility. Radio Frequency Applications 10 L-58 Inputs SpacecraftÐMissile System 9 L-58 Prelaunch Safety Package Table 8-2. MDA Program Documents (MSPSP) Nominal PMAÐPreliminary Mission 11 L-54/L-39 Due weeks Analysis Table 8-3 Ð or + Requirements/Comments Description reference launch Mission Operational and 12, 13 L-52 Mission Specification (Initial) 4 L-84 Support Requirements for Coupled Dynamic Loads 6 L-68 Spacecraft Analysis Payload Processing 14 L-52 Spacecraft-to-Blockhouse 29 L-50, L-24 Requirements Document Inputs Wiring Diagram (Prel/Final) Spacecraft-to-Blockhouse 29 L-40 Preliminary Mission Analysis 11 L-44 Wiring Diagram Review (PMA) DTOÐDetailed Test Objectives 17 L-39 Payload Processing 14 L-39 Requirements Requirements Document Launch Window (Initial/Final) 16 L-39, L-4 Spacecraft Compatibility 18 L-36, L-17 Vehicle Launch Insignia 15 L-39 Drawing Spacecraft Launch Site Test 19 L-34 Detailed Test Objective (DTO) 17 L-28 Plan Spacecraft-Fairing Clearance 18 L-27 Spacecraft Compatibility 18 L-29 Drawing Drawing Comments Launch Site Procedures Ð L-15 Combined Spacecraft/Third- 22 L-54/L-20 Integrated Countdown Schedule Ð L-15 Stage Nutation Time Constant Nutation Control System 23 L-15 and Mass Properties Statement Analysis (Initial/Final) Spacecraft Separation Analysis 25 L-12 Spacecraft Integrated Test 21 L-20 Launch Operations Plan 26 L-4 Procedures VIM—Vehicle Information 27 L-3 Spacecraft Launch Site Test 20 L-18 Memorandum Procedures M029, T030-,9/22/95 , 09 :21 AM Spacecraft Environments and 5 L-18 Loads Test Report Mission Operational and 12 L-12 Support Requirements Postlaunch Orbit Confirmation 28 L+1 Data

8-4 Table 8-3. Required Documents Item Responsibility 1. Feasibility Study (Optional) A feasibility study may be necessary to define the launch vehicle's capabilities for a specific mission or to establish the overall feasibility of using the vehicle for performing the required mission. Typical items that may necessitate a feasibility study are (1) a new flight plan with unusual launch azimuth or orbital requirements; (2) a precise accuracy requirement or a performance requirement greater than that available with the standard vehicle; and (3) spacecraft that impose uncertainties with regard to vehicle MDA stability. Specific tasks, schedules, and responsibilities are defined before study initiation, and a final report is prepared at the conclusion of the study. 2. Spacecraft Questionnaire The Spacecraft Questionnaire (Table 8-4) is the first step in the process and is designed to provide the initial definition of spacecraft requirements, interface details, launch site facilities, and preliminary safety data to Delta's various agencies. It contains a set of questions whose answers define the requirements and interfaces as they are known at the time of preparation. The questionnaire is required not later than 2 years prior to launch. Spacecraft A definitive response to some questions may not be possible because many items are defined at a later agency date. Of particular interest are answers that specify requirements in conflict with constraints specified herein. Normally this document would not be kept current; it will be used to create the initial issue of the Mission Specification (Item 4) and in support of our Federal Aviation Administration (FAA) launch permit. The specified items are typical of the data required for Delta II missions. The spacecraft agency is encouraged to include other pertinent information regarding mission requirements or constraints. 3 Spacecraft Mathematical Model for Dynamic Analysis A spacecraft mathematical model is required for use in a coupled loads analysis. Acceptable forms include (1) a discrete math model with associated mass and stiffness matrices or (2) a constrained normal mode model with modal mass and stiffness and the appropriate transformation matrices to Spacecraft recover internal responses. Required model information such as specific format, degree of freedom agency requirements, and other necessary information will be supplied. 4. Mission Specification The MDA Mission Specification functions as the Delta launch vehicle interface control document and describes all mission-specific requirements. It contains the spacecraft description, spacecraft-to-blockhouse wiring diagram, compatibility drawing, targeting criteria, special spacecraft requirements affecting the standard launch vehicle, description of the mission-specific vehicle, a MDA description of special AGE and facilities MDA is required to furnish, etc. The document is provided to (input spacecraft agencies for review and concurrence and is revised as required. The initial issue is based required upon data provided in the Spacecraft Questionnaire and is provided approximately 84 weeks before from user) launch. Subsequent issues are published as requirements and data become available. The mission peculiar requirements documented in the Mission Specification along with the standard interfaces presented in this manual define the spacecraft to launch vehicle interface. 5. Spacecraft Environmental Test Documents The environmental test plan documents the spacecraft contractor's approach for qualification and acceptance (preflight screening) tests. It is intended to provide general test philosophy and an overview of the system-level environmental testing to be performed to demonstrate adequacy of the spacecraft for flight (e.g., static loads, vibration, acoustics, shock). The test plan should include test objectives, test specimen configuration, general test methods, and a schedule. It should not include detailed test procedures. Spacecraft agency Following the system-level structural loads and dynamic environment testing, test reports documenting the results shall be provided to MDA. These reports should summarize the testing performed to verify the adequacy of spacecraft structure for the flight loads. For structural systems not verified by test, a structural loads analysis report documenting the analyses performed and resulting margins of safety should be provided to MDA. 6. Coupled Dynamic Loads Analysis A coupled dynamic loads analysis is performed in order to define flight loads to major vehicle and spacecraft structure. The liftoff event, which generally causes the most severe lateral loads in the spacecraft, and the period of transonic flight and maximum dynamic pressure, causing the greatest MDA relative deflections between spacecraft and fairing, are generally included in this analysis. Output for (input each flight event includes tables of maximum acceleration at selected nodes of the spacecraft model as required well as a summary of maximum interface loads. Worst-case spacecraft-fairing dynamic relative from user, deflections are included. Close coordination between the user and the Delta Program Office is essential Item 3) in order to decide on the output format and the actual work schedule for the analysis.

8-5 Table 8-3. Required Documents (Continued) Item Responsibility 7. Electrical Wiring Requirements The wiring requirements for the spacecraft to the blockhouse and the payload processing facilities are needed as early as possible. Section 5 lists the Delta capabilities and outlines the necessary details to be supplied. MDA will provide a spacecraft-to-blockhouse wiring diagram based on the spacecraft requirements. It will define the hardware interface from the spacecraft to the blockhouse for control and Spacecraft monitoring of spacecraft functions after spacecraft installation in the launch vehicle. Close attention to agency the documentation schedule is required so that production checkout of the launch vehicle includes all of the mission-specific wiring. Any requirements for the payload processing facilities are to be furnished with the blockhouse information. 8. Fairing Requirements Early spacecraft fairing requirements should be addressed in the questionnaire and updated in the Mission Specification. Final spacecraft requirements are needed to support the mission-specific fairing Spacecraft modifications during production. Any in-flight requirements, ground requirements, critical spacecraft agency surfaces, surface sensitivities, mechanical attachments, RF transparent windows, and internal temperatures on the ground and in flight must be provided. 9. Missile System Prelaunch Safety Package (MSPSP) (Refer to EWR 127-1 for specific spacecraft safety regulations.) To obtain approval to use the launch site facilities and resources and for launch, a MSPSP must be prepared and submitted to the Delta Program Office. The MSPSP includes a description of each hazardous system (with drawings, schematics, and assembly and handling procedures, as well as any other information that will aid in appraising the respective systems) and evidence of compliance with Spacecraft the safety requirements of each hazardous system. The major categories of hazardous systems are agency ordnance devices, radioactive material, propellants, pressurized systems, toxic materials and cryogenics, and RF radiation. The specific data required and suggested formats are discussed in Section 2 of EWR 127-1. MDA will provide this information to the appropriate government safety offices for their approval. 10. Radio Frequency Applications The spacecraft agency is required to specify the RF transmitted by the spacecraft during ground processing and launch intervals. A RF data sheet specifying individual frequencies will be provided. Names and qualifications are required covering spacecraft user personnel who will operate spacecraft Spacecraft RF systems. Transmission frequency bandwidths, frequencies, radiated durations, wattage etc., will be agency provided. MDA will provide these data to the appropriate range/government agencies for approval. 11. Preliminary Mission Analysis (PMA) This analysis is normally the first step in the mission planning process. It uses the best available mission requirements (spacecraft weight, orbit requirements, tracking requirements, etc.) and is primarily intended to uncover and resolve any unusual problems inherent in accomplishing the mission objectives. Specifically, information pertaining to vehicle environment, performance capability, MDA sequencing, and orbit dispersion is presented. Parametric performance and accuracy data are usually (input provided to assist the user in selection of final mission orbit requirements. The orbit dispersion data are required presented in the form of variations of the critical orbit parameters as functions of probability level. A from user) covariance matrix and a trajectory printout are also included. The mission requirements and parameter ranges of interest for parametric studies are due as early as possible but in no case later than 54 weeks before launch. Comments to the PMA are needed no later than launch minus 39 weeks for start of the DTO (Item 17). 12. Mission Operational and Support Requirements To obtain unique range and network support, the spacecraft agency must define any range or network requirements appropriate to its mission and then submit them to MDA. Spacecraft agency operational Spacecraft configuration, communication, tracking, and data flow requirements are required to support document agency preparation and arrange required range support. 13. Program Requirements Documents (PRD) To obtain range and network support, a spacecraft PRD must be prepared. This document consists of a MDA set of preprinted standard forms (with associated instructions) that must be completed. The spacecraft (input agency will complete all forms appropriate to its mission and then submit them to MDA. MDA will required compile, review, provide comments, and, upon comment resolution, forward the spacecraft PRD to the from user) appropriate support agency for formal acceptance.

8-6 Table 8-3. Required Documents (Continued) Item Responsibility 14. Payload Processing Requirements Documents (PPRD) The PPRD is prepared if commercial facilities are to be used for spacecraft processing. The spacecraft agency is required to provide data on all spacecraft activities to be performed at the commercial facility. This includes detailed information of all facilities, services, and support requested by MDA to be pro- Spacecraft vided by the commercial facility. Spacecraft hazardous systems descriptions shall include drawings, agency schematics, summary test data, and any other available data that will aid in appraising the respective hazardous system. The commercial facility will accept spacecraft ground operations plans and/or MSPSP data for the PPRD. 15. Launch Vehicle Insignia The spacecraft customer is entitled to have a mission-specific insignia placed on the launch vehicle. The customer will submit the proposed design to MDA not later than 9 months before launch for review Spacecraft and approval. Following approval, MDA will have the flight insignia prepared and placed on the launch agency vehicle. The maximum size of the insignia is 2.4 by 2.4 m (8 by 8 ft). The insignia is placed on the uprange side of the launch vehicle. 16. Launch Window The spacecraft agency is required to specify the maximum launch window for any given day. Specifically the window opening time (to the nearest minute) and the window closing time (to the Spacecraft nearest minute) are to be specified. This final window data should extend for at least 2 weeks beyond agency the scheduled launch date. Liftoff is targeted to the specified window opening. 17. Detailed Test Objectives (DTO) Report MDA will issue a DTO trajectory report that provides the mission reference trajectory. The DTO contains a description of the flight objectives, the nominal trajectory printout, a sequence of events, vehicle attitude rates, spacecraft and vehicle tracking data, and other pertinent information. The trajectory is MDA used to develop mission targeting constants and represents the flight trajectory. The DTO will be (input required available at launch minus 28 weeks. from user) 18. Spacecraft Drawings Spacecraft configuration drawings are required as early as possible. The drawings should show nominal and worst-case (maximum tolerance) dimensions for the MDA-prepared compatibility drawing, clearance analysis, fairing compatibility, and other interface details. Preliminary drawings are desired Spacecraft with the Spacecraft Questionnaire but no later than 78 weeks prior to launch. The drawings should be agency 0.20 scale for 9.5- and 10-ft diameter fairings. The drawings should be supplied on vellum or mylar and rolled. The capability exists for CAD drawing use. Details should be worked through the Delta Program Office. MDA will prepare and release the spacecraft compatibility drawing that will become part of the Mission Specification. This is a working drawing that identifies spacecraft to launch vehicle interfaces. It defines MDA electrical interfaces; mechanical interfaces, including spacecraft-to-PAF separation plane, separation springs and spring seats, and separation switch pads; definition of stay-out envelopes, both internal and external to the PAF; definition of stay-out envelopes within the fairing; and location and mechanical activation of spring seats. The user agency reviews the drawing and provides comments, and upon comment resolution and incorporation of the final spacecraft drawings, the compatibility drawing is formally accepted as a controlled interface between MDA and the spacecraft agency. In addition, MDA will provide a worst-case spacecraft-fairing clearance drawing. 19. Spacecraft Launch Site Test Plan To provide all agencies with a detailed understanding of the launch site activities and operations planned for a particular mission, the spacecraft agency is required to prepare a Launch Site Test Plan. Spacecraft The plan is intended to describe all aspects of the program while at the launch site. A suggested format agency is shown in Table 8-5. 20. Spacecraft Launch Site Test Procedures Operating procedures must be prepared for all operations that are accomplished at the launch site. For Spacecraft those operations that are hazardous in nature (either to equipment or to personnel), special instructions agency must be followed in preparing the procedures. Refer to Section 9.

8-7 Table 8-3. Required Documents (Continued) Item Responsibility 21. Spacecraft Integrated Test Procedure Inputs On each mission, MDA prepares launch site procedures for various operations that involve the spacecraft after it is mated with the Delta upper stage. Included are requirements for operations such as spacecraft weighing, spacecraft installation to third stage and into the handling can, spacecraft transportation to the launch complex, spacecraft hoisting into the white room, handling-can removal, Spacecraft spacecraft/third stage mating to launch vehicle, fairing installation, flight program verification test, and agency launch countdown. MDA requires inputs to these operations in the form of handling constraints, environmental constraints, personnel requirements, equipment requirements, etc. Of particular interest are spacecraft tasks/requirements during the final week before launch. (Refer to Section 6 for schedule constraints.) 22. Spacecraft Mass Properties Statement and Nutation Time Constants The combined spacecraft/third-stage nutation time constant for preburn and postburn conditions is required before launch so that the effects of energy dissipation relative to spacecraft separation, coning buildup, and clearance during separation can be evaluated. The data from the spacecraft mass properties report is used in spin rocket configuration, orbit error, control, performance, and separation Spacecraft analyses. It represents the best current estimate of final spacecraft mass properties. The data should agency include any changes in mass properties while the spacecraft is attached to the Delta vehicle. Values quoted should include nominal and 3σ uncertainties for mass, centers of gravity, moments of inertia, products of inertia, and principal axis misalignment and Delta upper stage mass properties provided in Section 4.2. 23. Nutation Control System Analysis Memorandum A Nutation Control System (NCS) analysis is performed to verify that the system is capable of controlling the third-stage coning motion induced by the dynamic-coupled instability. The NCS is activated at third-stage ignition and remains active throughout the burn and coast until the start of NCS blowdown. The principal inputs required for the analysis are the spacecraft mass properties and MDA nutation time constants from Item 22 and the third-stage mass properties. The analysis outputs include spacecraft/third-stage rates and angular momentum pointing prior to spacecraft separation, third-stage velocity loss and pointing error (used in orbit dispersion analysis), and NCS propellant usage. 24. RF Compatibility Analysis A radio frequency interference (RFI) analysis is performed to verify that spacecraft RF sources are compatible with the launch vehicle telemetry and tracking beacon frequencies. Spacecraft frequencies defined in the mission specification are analyzed using a frequency compatibility software program. The MDA program provides a listing of all intermodulation products which are then checked for image frequencies and intermodulation product interference. 25. Spacecraft/Launch Vehicle Separation Memorandum An analysis is performed to verify that there is adequate clearance and separation distance between the MDA spacecraft and expended PAF/third stage. The principal parameters, including data from Item 22, that (input define the separation are the motor's residual thrust, half-cone angle, and spin rate. For two-stage required missions this analysis verifies adequate clearance between the spacecraft and second stage during from user) separation and second-stage post-separation maneuvers. 26. Launch Operations Plan (LOP) This plan is developed to define top-level requirements that flow down into detailed range requirements. The plan contains the launch operations configuration, which identifies data and communication connectivity with all required support facilities. The plan also identifies organizational roles and MDA responsibilities, the mission control team and its roles and responsibilities, mission rules supporting conduct of the launch operation, and go/no-go criteria. 27. Vehicle Information Memorandum (VIM) MDA is required to provide a Vehicle Information Message to the US Space Command 15 calendar days prior to launch. The spacecraft agency will provide to MDA the appropriate spacecraft on-orbit data required for this VIM. Data required are spacecraft on-orbit descriptions, description of pieces and MDA debris separated from the spacecraft, the orbital parameters for each piece of debris, S/C spin rates, and orbital parameter information for each different orbit through final orbit. MDA will incorporate these data into the overall VIM and transmit to the appropriate US government agency. 28. Postlaunch Orbit Confirmation Data To reconstruct Delta performance, orbit data at burnout (stage II or III) are required from the spacecraft agency. The spacecraft agency should provide orbit conditions at the burnout epoch based on Spacecraft spacecraft tracking data prior to any orbit correction maneuvers. A complete set of orbital elements and agency associated estimates of 3σ accuracy are required (see Table 8-6). 29. Spacecraft-to-Blockhouse Wiring Diagram MDA will provide, for inclusion into the Mission Specification, a spacecraft-to-blockhouse wiring diagram based on the spacecraft requirements. It will define the hardware interface from the spacecraft to the MDA blockhouse for control and monitoring of spacecraft functions after spacecraft installation in the launch vehicle. M029, T031, 9/22/95, 9:21 AM 8-8 Table 8-4. Spacecraft Questionnaire (Download Table) 1 Spacecraft Characteristics 1.1 Mission Objectives 1.2 Size and Space Envelope 1.2.1 Dimensioned Drawings/CAD Model (Launch Configuration) 1.2.2 Protuberances Within 50.8 mm/2.0 in. of Allowable Fairing Envelope Below Separation Plane (Identify Component and Location) 1.2.3 Appendages Below Separation Plane (Identify Component and Location) 1.3 Orbit Configuration 1.4 Mass Properties and Dynamic Data (Launch Configuration/Separation Configuration Including Tolerances) 1.4.1 Weight (Maximum or Nominal) 1.4.2 CG Location 1.4.3 Moments of Inertia About Spacecraft CG in Body Axes (Ixx, Iyy, Izz) 1.4.4 Products of Inertia (Ixy, Ixz, Iyz) 1.4.5 Spacecraft Coordinate System 1.4.6 Fundamental Frequencies (Thrust Axis/Lateral Axis) 1.4.7 Are All Significant Vibration Modes Above 35 Hz in Thrust and 15 Hz (12 Hz for two stage) in Lateral Axes? 1.4.8 Description of Spacecraft Dynamic Model Mass Matrix Stiffness Matrix Response Recovery Matrix 1.4.9 Combined Spacecraft-Third Stage Nutation Time Constant for Ignition and Burnout Conditions (If Applicable) 1.4.10 Time Constant and Description of Spacecraft Energy Dissipation Sources and Locations (i.e., Hydrazine Fill Factor, Passive Nutation Dampers, Flexible Antennae, etc.) 1.5 Spacecraft Hazardous Systems 1.5.1 Apogee Motor (Solid or Liquid) 1.5.2 Attitude Control System 1.5.3 Hydrazine (Quantity, Spec, etc.) 1.5.4 Other Hazardous Fluids (Quantity, Spec, etc.) 1.5.5 High Pressure Gas (Quantity, Spec, etc.) 1.5.6 Radioactive Devices 1.5.7 Can Spacecraft Produce Nonionizing Radiation at Hazardous Levels? 1.5.8 Do Pressure Vessels Conform to Safety Requirements of Delta Payload Planners Guide Section 9? 1.5.9 Location Where Pressure Vessels Are Loaded and Pressurized 1.5.10 Other Hazardous Systems 1.6 Electro-Explosive Devices (EED) 1.6.1 Category A EEDs (Function, Type, Part Number, When Installed, When Connected) 1.6.2 Are Electrostatic Sensitivity Data Available on Category A EEDs? List References 1.6.3 Category B EEDs (Function, Type, Part Number, When Installed, When Connected) 1.6.4 Do Shielding Caps Comply With Safety Requirements? 1.6.5 Are RF Susceptibility Data Available? List References 1.7 RF Systems 1.7.1 System 1.7.2 Frequency (MHz) 1.7.3 Maximum Power (EIRP) (dBm) 1.7.4 Average Power (W) 1.7.5 Type of Transmitter 1.7.6 Antenna Gain (dB) 1.7.7 Antenna Location 1.7.8 Distance at Which RF Radiation Flux Density Equals 1 mW/cm2 1.7.9 When is RF Transmitter Operated? 1.7.10 RF Checkout Requirements (Location and Duration, to What Facility, Support Requirements, etc.) 1.8 Contamination-Sensitive Surfaces 1.9 Spacecraft Volume (Ventable and Non-Ventable) 1.10 Spacecraft Systems Activated Prior to Separation from Delta 2 Mission Requirements and Restraints 2.1 Number of Launches 2.2 Frequency of Launches 2.3 Desired Transfer Orbit 2.3.1 Apogee (Integrated) 2.3.2 Perigee (Integrated) 2.3.3 Inclination 2.3.4 Argument of Perigee at Insertion 2.3.5 Other

8-9 Table 8-4. Spacecraft Questionnaire (Continued) 2.4 Launch Window Restraints and Duration 2.4.1 Sun Angle 2.4.2 Eclipse 2.4.3 Ascending Node 2.4.4 Inclination 2.4.5 Window Durations (Over a Year’s Span) 2.4.6 Other 2.5 Separation Requirements (Including Tolerances) 2.5.1 Position 2.5.2 Attitude 2.5.3 Sequence and Timing 2.5.4 Tipoff and Coning 2.5.5 Spin Rate at Separation 2.5.6 Other 2.6 Special Trajectory Requirements 2.6.1 Thermal Maneuvers 2.6.2 T/M Maneuvers 2.6.3 Free Molecular Heating Restraints 3 Spacecraft-to-Launch Vehicle Interface Requirements 3.1 PAF Desired (3712A, 6915, etc.) 3.2 Interface Connector and Umbilicals 3.2.1 Type and Part Number 3.2.2 Location 3.2.3 Orientation 3.3 Electrical BondingÐDoes The Spacecraft Comply With the Electrical Bonding Requirements of Section 4? 3.4 Spacecraft-to-Blockhouse Wiring Requirements 3.4.1 Number of Wires Required 3.4.2 Pin Assignments in the Spacecraft Umbilical Connector(s) 3.4.3 Purpose and Nomenclature of Each Wire Including Voltage, Current, Polarity Requirements, and Maximum Resistance 3.4.4 Shielding Requirements 3.4.5 Voltage of the Spacecraft Battery and Polarity of the Battery Ground 3.5 Does Spacecraft Require Discrete Signals from Delta? 3.6 Separation Switch Pads 3.7 Separation Switches 3.8 Separation Springs 3.9 Fairing Required (9.5 or 10 ft) 3.10 Access Doors in Fairing (Size, Location, Purpose) 3.11 RF Window Requirements (Size, Location, Purpose) 3.12 Fairing Environmental Requirements 3.12.1 Spacecraft In-Flight Requirements 3.12.2 Spacecraft Ground Requirements (Fairing Installed) 3.12.3 Critical Surfaces (i.e., Type, Size, Location) 3.12.4 Surface Sensitivity (e.g., Susceptibility to Propellants, Gases and Exhaust Products, and Other Contaminants 3.12.5 Purges Required 4 Spacecraft Processing Facility Requirements 4.1 Processing Facility Preference and Priority 4.2 List The Hazardous Processing Facilities the Spacecraft Project Desires to Use 4.3 What Are The Expected Dwell Times the Spacecraft Project Would Spend in the Payload Processing Facilities? 4.4 Spacecraft Environmental Requirements 4.4.1 Cleanliness Requirements 4.4.2 Temperature Requirements 4.4.3 Humidity Requirements 4.4.4 Do Spacecraft Contamination Requirements Conform With Capabilities of Existing Facilities? 4.5 What Are the Spacecraft and Ground Equipment Space Requirements? 4.6 What Are the Facility Crane Requirements? 4.7 What Are the Facility Electrical Requirements? 4.8 What Are the Security Requirements? 4.9 Is a Multishift Operation Planned? 4.10 List the Support Items the Spacecraft Project Needs from NASA, USAF, or Commercial Providers to Support the Processing of Spacecraft. Are There any Unique Support Items?

8-10 Table 8-4. Spacecraft Questionnaire (Continued) 4.11 Special MDA Handling Requirements 4.11.1 In Payload Processing Facility (PPF) 4.11.2 In Handling Can 4.11.3 On Stand 4.11.4 In Blockhouse 4.12 Special MDA-Supplied AGE or Facilities 4.13 Spacecraft-To-Blockhouse Signal Conditioning Requirements 5 Spacecraft Development and Test Programs 5.1 Test Schedule at Launch Site 5.1.1 Operations Flow Chart (Flow Chart Should Be a Detailed Sequence of Operations Referencing Days, Shifts, and Location) 5.2 Spacecraft Development and Test Schedules 5.2.1 Flow Chart and Test Schedule 5.2.2 Is a Test PAF Required? When? 5.2.3 Is Clamp Band Ordnance Required? When? 5.3 Special Test Requirements 5.3.1 Spacecraft Spin Balancing 5.3.2 Other 6 Identify Any Additional Spacecraft or Mission Requirements that Exceed the Capabilities or Violate the Constraints Defined in the Payload Planner’s Guide

8-11 Table 8-5. Typical Spacecraft Launch Site Test Plan (Download Table) 1 General 1.1 Plan Organization 1.2 Plan Scope 1.3 Applicable Documents 1.4 Spacecraft Hazardous Systems Summary 2 Prelaunch/Launch Test Operations Summary 2.1 Schedule 2.2 Layout of Equipment (Each Facility) (Including Test Equipment) 2.3 Description of Event at Launch Site 2.3.1 Spacecraft Delivery Operations 2.3.1.1 Spacecraft Removal and Transport to Spacecraft Processing Facility 2.3.1.2 Handling and Transport of Miscellaneous Items (Ordnance, Motors, Batteries, Test Equipment, Handling and Transportation Equipment) 2.3.2 Payload Processing Facility Operations 2.3.2.1 Spacecraft Receiving Inspection 2.3.2.2 Battery Inspection 2.3.2.3 RCS Leak Test 2.3.2.4 Battery Installation 2.3.2.5 Battery Charging 2.3.2.6 Spacecraft Validation 2.3.2.7 Solar Array Validation 2.3.2.8 Spacecraft/Data Network Compatibility Test Operations 2.3.2.9 Spacecraft Readiness Review 2.3.2.10 Preparation for Transport, and Transport to HPF 2.3.3 Solid Fuel Storage Area 2.3.3.1 AKM Receiving, Preparation, and X-Ray 2.3.3.2 S&A Device Receiving, Inspection, and Electrical Test 2.3.3.3 Igniter Receiving and Test 2.3.3.4 AKM/S&A Assembly and Leak Test 2.3.4 HPF 2.3.4.1 Spacecraft Receiving Inspection 2.3.4.2 Preparation for AKM Installation 2.3.4.3 Mate AKM to Spacecraft 2.3.4.4 Spacecraft Weighing (Include Configuration Sketch and Approximate Weights of Handling Equipment) 2.3.4.5 Spacecraft/Third-Stage Mating 2.3.4.6 Preparation for Transport Installation Into Handling Can 2.3.4.7 Transport to Launch Complex 2.3.5 Launch Complex Operations 2.3.5.1 Spacecraft Hoisting and Removal of Handling Can 2.3.5.2 Spacecraft Mate to Launch Vehicle 2.3.5.3 Hydrazine Leak Test 2.3.5.4 TT&C Checkout 2.3.5.5 Preflight Preparations 2.3.5.6 Fairing Installation 2.3.5.7 Launch Countdown 2.4 Launch/Hold Criteria 2.5 Environmental Requirement for Facilities During Transport 3 Test Facility Activation 3.1 Activation Schedule 3.2 Logistics Requirements 3.3 Equipment Handling 3.3.1 Receiving 3.3.2 Installation 3.3.3 Validation 3.3.4 Calibration 3.4 Maintenance 3.4.1 Spacecraft 3.4.2 Launch Critical Mechanical AGE and Electrical AGE 4 Administration 4.1 Test Operations—Organizational Relationships and Interfaces (Personnel Accommodations, Communications) 5 Security Provisions for Hardware 6 Special Range Support Requirements 6.1 Real-Time Tracking Data Relay Requirements 6.2 Voice Communications 6.3 Mission Control Operations M029: T034.1 4/16/93 3:01 PM

8-12 Table 8-6. Data Required for Orbit Parameter Statement (Download Table) 1. Epoch: Third-stage burnout . . . 2. Position and velocity components (X, Y, Z, X, Y, Z) in equatorial inertial Cartesian coordinates.* Specify mean-of-date or true-of-date, etc.

3. Keplerian elements* at the above epoch: Semimajor axis, a Eccentricity, e Inclination, i Argument of perigee, ω Mean anomaly, M Right ascension of ascending node, Ω

4. Polar elements* at the above epoch: Inertial velocity, V Inertial flight path angle, γ1 Inertial flight path angle, γ2 Radius, R Geocentric latitude, ρ Longitude, µ

5. Estimated accuracies of elements and a discussion of quality of tracking data and difficulties such as reorientation maneuvers within 6 hours of separation, etc.

6. Constants used: Gravitational constant, µ Equatorial radius, RE J2 or Earth model assumed

7. Estimate of spacecraft attitude and coning angle at separation (if available). *Note: At least one set of orbit elements in Items 2, 3, or 4 is required. M029:T033 4/16/93 2:16 PM

8-13 H30573.5

Weeks 100 9080 7060 5040 3020 10 0 Agency Milestones

Spacecraft Spacecraft Questionnaire L-104 L-90 Launch Spacecraft Spacecraft Mathematical Model L-84 Spacecraft Spacecraft Environment Test Document MDA Mission Specification L-84 Initial L-78 Initial Spacecraft Spacecraft Drawings L-44 Final Spacecraft Mission Specification Comments L-80 MDA Coupled Dynamic Loads Analysis L-68 Spacecraft Fairing Requirements L-68 Spacecraft Electrical Wiring Requirements L-80 Spacecraft Spacecraft Missile System Prelaunch Safety L-58 PackagePackage (MSPSP) (MSPSP) Spacecraft Radio Frequency Application (RFA) L-58 Spacecraft Preliminary Mission Analysis (PMA) Requirements L-54 Spacecraft Payload Processing Requirements Doc (PPRD) Input L-52 Spacecraft Mission Operations and Support Requirements L-52 Final MDA Spacecraft-to-Blockhouse Wiring Diagram Preliminary L-50 L-24 MDA Preliminary Mission Analysis L-44 Spacecraft Spacecraft-to-Blockhouse Wiring Diagram Comments L-40 Spacecraft Launch Vehicle Insignia L-39 Final Spacecraft Launch Window L-39 Initial L-4 Spacecraft Detailed Test Objective (DTO) Requirements L-39 MDA Payload Processing Requirements Document L-39 MDA Spacecraft Compatibility Drawing L-36 L-17 Final Spacecraft Spacecraft Launch Site Test Plan L-34 Spacecraft Spacecraft Compatibility Drawing Comments L-29 MDA Detailed Test Objective L-28 MDA Spacecraft Fairing Clearance Drawing L-27 MDA Program Requirements Document L-26 Spacecraft Combined Spacecraft/Third-Stage Nutation Time L-54 Initial L-20 Final ConstantConstant & & Mass Mass Properties Properties Statement Statement Spacecraft Spacecraft Integrated Test Procedure Input L-20 Spacecraft Spacecraft Launch Site Procedures L-18 Spacecraft Spacecraft Environments and Loads Test Report L-18 MDA Launch Site Procedures L-15 MDA Integrated Countdown Schedule L-15 MDA Nutation Control System Analysis L-15 MDA RF Compatibility Study Results L-12 MDA Spacecraft Separation Analysis L-12 Final MDA Launch Operations Plan L-12 L-4 MDA Vehicle Information Memo (VIM) L-3 Spacecraft Postlaunch Orbit Confirm. Data (Orbital Tracking Data) L+1 Day MDA Postlaunch Flight Report L+8

Launch

Figure 8-4. Typical Integration Planning Schedule 8-14 H30574.1

Launch

Weeks Pre Post 60 50 40 30 20 10 0 +10 +20

-54-52 -39 Spacecraft Agency Inputs DTO Mission Requirements Preliminary Mission Requirements -44 PMA AAAAAAAAAA-28 DTO Mission Definition Preliminary OperationalAAAAAAAAAA Configuration Requirements -30 Days Spacecraft PRD Inputs Launch Operation Plan AAAAAAAAAA

-26 PRD (Update As Required) PI (If Required) Range Support Requirements AAAAAA AAAAAA-12 Mission Support Request NASA Support Requirements AAAAAA Figure 8-5. Launch Operational Configuration Development

8-15 Table 8-7. Spacecraft Checklist (Download Table) 1. General (10) Data lines (from/to where) ______A. Transportation of spacecraft elements/GSE to (11) Type (wideband/narrowband) ______processing facility H. Services general (1) Mode of transportation: (1) Gases (2) Arriving at ______(gate, skid a. Specification ______strip) Procured by user? ______KSC?_____ (date)______b. Quantity ______B. Data handling c. Sampling: (yes) ______(no) ______(1) Send data to (name and address) (2) Photographs/video _____ (quantity/B&W/color) (2) Time needed (real time versus after the fact) (3) Janitorial (yes) ______(no) ______C. Training and medical examinations for (4) Reproduction services (yes) _____ (no) ______crane operators I. Security (yes) ______(no) ______D. Radiation data (1) Safes ______(number/type) (1) Ionizing radiation materials J. Storage ______(size area) (2) Nonionizing radiation materials/systems ______environment 2. Spacecraft Processing Facility (for nonhazardous K. ______work) L. Spacecraft PPF activities calendar A. Does payload require a cleanroom? (1) Assembly and testing ______(yes) ____ (no) ____ (2) Hazardous operations (1) Class of cleanroom required: a. Initial turn-on of a high-power RF system (2) Special sampling techniques: ______B. Area required: b. Category B ordnance installation _____ (1) For spacecraft ______sq ft c. Initial pressurization ______(2) For ground station ______sq ft d. Other ______(3) For office space ______sq ft M. Transportation of payloads/GSE from PPF to HPF (4) For other GSE ______sq ft (1) Will spacecraft agency supply transportation (5) For storage ______sq ft canister? ______C. Largest door size: If no explain ______(1) For spacecraft/GSE ______(2) Equipment support, e.g., mobile crane, flatbed (high) ______(wide) ______(2) For ground station: (3) Weather forecast (yes) ______(no) ______D. Material handling equipment: (4) Security escort (yes) ______(no) ______(1) Cranes (5) Other ______a. Capacity: 3. Hazardous Processing Facility b. Minimum hook height: A. Does spacecraft require a cleanroom? c. Travel: ______(yes) _____ (no) (2) Other ______(1) Class of cleanroom required: E. Environmental controls for spacecraft/ground (2) Special sampling techniques: (e.g., station hydrocarbon monitoring) (1) Temperature/humidity and tolerance limits: B. Area required: (2) Frequency of monitoring (1) For spacecraft ______sq ft (3) Downtime allowable in the event of a system (2) For GSE ______sq ft failure ______C. Largest door size: (4) Is a backup (portable) air-conditioning system (1) For payload ______high ______wide required? (yes) ______(no) ______(2) For GSE ______high ______wide (5) ______D. Material handling equipment F. Electrical power for payload and ground station (1) Cranes (1) kVA required: a. Capacity: (2) Any special requirements such as clean/quiet b. Hook height: power, or special phasing? c. Travel ______Explain ______(2) Other ______(3) Backup power (diesel generator) E. Environmental controls spacecraft/GSE a. Continuous: (1) Temperature/humidity and tolerance limits: b. During critical tests: (2) Frequency of monitoring ______G. Communications (list) (3) Down-time allowable in the event of a system (1) Administrative telephone failure ______(2) Commercial telephone ______(4) Is a backup (portable) system required? (3) Commercial data phones ______(yes) _____ (no) ______(4) TWX machines ______(5) Other ______(5) Operational intercom system ______F. Power for spacecraft and GSE (6) Closed circuit television ______(1) kVA required: (7) Countdown clocks ______G. Communications (list) (8) Timing ______(1) Administrative telephone ______(9) Antennas ______(2) Commercial telephone ______

M029, T035, 7/15/97, 10:25 AM

8-16 Table 8-7. Spacecraft Checklist (Continued) (3) Commercial data phones ______(4) Hydrocarbon monitoring required ______(4) TWX machines ______(5) Frequency of monitoring ______(5) Operational intercom system ______(6) Down-time allowable in the event of a system (6) Closed circuit television ______failure ______(7) Countdown clocks ______(7) Other ______(8) Timing ______B. Power for payload and GSE (9) Antennas ______(1) kVA required ______(10) Data lines (from/to where) ______(2) Any special requirements such as clean/quiet H. Services general power/phasing? (1) Gases Explain: ______a. Specification ______(3) Backup power (diesel generator) Procured by user? _____ KSC? ______a. Continuous: ______b. Quantity ______b. During critical tests: ______c. Sampling? (yes) _____ (no) ______C. Communications (list) (2) Photographs/video ___ (quantity/B&W/color) (1) Operational intercom system ______(3) Janitorial (yes) ______(no) ______(2) Closed circuit television ______(4) Reproduction services (yes) ____ (no) _____ (3) Countdown clocks ______I. Security (yes) ______(no) ______(4) Timing ______J. Storage ______(size area) (5) Antennas ______(environment)______(6) Data lines (from/to where) ______K. Other _____ D. Services general L. Spacecraft HPF activities calendar (1) Gases (1) Assembly and testing ______a. Specification ______(2) Hazardous operations Procured by user? _____ KSC? ______a. Category A ordnance installation ______b. Quantity ______b. Fuel loading ______c. Sampling? (yes) ______(no) ______c. Mating operations (hoisting) (2) Photographs ______(quantity/B&W/color) M. Transportation of payloads to LC17 E. Security (yes) _____ (no) ______(1) Equipment support, e.g., mobile crane, flatbed F. Other ______G. Stand-alone testing (does not include tests involving (2) Weather forecast (yes) _____ (no) ______the Delta II vehicle) (3) Security escort (yes) _____ (no) ______(1) Tests required ______(4) Other ______(e.g., RF system checkout, encrypter checkout) 4. Launch Complex White Room (MST) (2) Communications required for ______A. Environmental controls payload/GSE (e.g., antennas, data lines) (1) Temperature/humidity and tolerance limits __ (3) Spacecraft servicing required ______(2) Any special requirements such as clean/quiet (e.g., cryogenics refill) power? Explain: ______(3) Backup power (diesel generator) a. Continuous: b. During critical tests: M029, T035.1, 9/20/95, 1:34 PM

8-17 Section 9 gram for operations conducted at NASA facilities. SAFETY In general, USAF and NASA safety regulations for This section discusses the safety regulations and prelaunch activities are equivalent. requirements that govern a payload to be launched Because many spacecraft operations are performed by or involve launch vehicle contractor by a Delta launch vehicle. Regulations and instruc- personnel, certain additional MDC safety require- tions that apply to spacecraft design and processing ments also apply. procedures are reviewed. The following documents specify the safety re- 9.1 SAFETY REQUIREMENTS quirements applicable to Delta II users: Delta II spacecraft prelaunch operations are con- A. EWR 127-1, Range Safety Requirements, 31 March 1995. ducted at Cape Canaveral Air Station (CCAS), B. KMI 1710.1G, Safety, Reliability, and Qual- Florida; Astrotech in Titusville, Florida; Kennedy ity Assurance Programs, 5 September 1991. Space Center (KSC), Florida; and Astrotech; Cali- C. Astrotech Space Operations Safety Standard fornia Spaceport, Vandenberg Air Force Base Operating Procedure (SOP), 1988 (VAFB), California, by arrangement with the ap- D. Astrotech Space Operations Safety Standard propriate agencies. The USAF is responsible for Operating Procedure at Vandenberg AFB, Septem- overall safety at CCAS and VAFB and has estab- ber 22, 1994. lished safety requirements accordingly. NASA The above regulations and instructions reference safety regulations govern operations at KSC, and other documents that further detail requirements for operations at Astrotech/California Spaceport are spacecraft design and operations. In addition, the conducted in accordance with Astrotech/California space wing safety regulations are supplemented by Spaceport safety policies. policy letters issued by the controlling organiza- Before a spacecraft moves onto USAF property, tions. Policy letters change, clarify, or add to the the spacecraft agency must provide the Space Wing existing regulations between formal revisions and (SW) safety office with certification that its system have the same effectivity as the regulation. The has been designed and tested in accordance with Space Wing safety organizations encourage pay- Space Wing safety requirements (EWR 127-1 Range load organizations to coordinate with them to gen- Safety Requirements). Safety of operations con- erate a tailored version of the 127-1 document ducted at NASA facilities on KSC and at VAFB specific to each program. This process can greatly must comply with KSC Management Instruction simplify the safety process at either range. MDA (KMI) 1710.1G. For operations at the Eastern provides coordination and assistance to the space- Range, both 45th Space Wing and NASA safety craft agency in this process. requirements are applicable. NASA/KSC at VAFB In addition to the above regulations, the space- and 30th SW have safety responsibility at Astrotech- craft agency must comply with MDA safety re- W, the NASA processing facilities, and at the SLC- quirements specified in the launch site safety plan 2W launch pad. The NASA safety requirements for operating at SLC-2. A specific spacecraft ad- and quality assurance (SR&QA) and protective dendum to this plan may be required for unique services directorate implement the KSC safety pro- spacecraft operations. 9-1 9.2 DOCUMENTATION REQUIREMENTS KMI 1710.1G from the outset of a program, utilize Both USAF and NASA require formal submittal them for design guidance, and submit the required of safety documentation containing detailed infor- data as early as possible. Document applicability is mation on all hazardous systems and associated determined by mission type and launch site as operations. The 30th and 45th Space Wings (30 SW shown in Table 9-1. & 45 SW) at the Western and Eastern Range require The safety document is submitted to the appro- preparation and submittal of a Missile System Pre- priate government agency or to MDA for commer- launch Safety Package (MSPSP). The content and cial missions for review and further distribution. format requirements of the document are found in Sufficient copies of the original and all revisions the Space Wing 127-1 Range Safety Requirements must be submitted by the originator to enable a and should be included in the tailoring process. review by all concerned agencies. The review pro- Data requirements for both ranges include design, cess usually requires several iterations until the test, and operational considerations. NASA require- system design and its intended use are considered to ments in almost every instance are covered by the be final and in compliance with all safety require- USAF requirements; however, the spacecraft agency ments. The flow of spacecraft safety information is can refer to KMI 1710.1G for details. dependent on the range to be used, the customer, A Ground Operations Plan must be submitted and contractual arrangements. Figure 9-1 illus- describing hazardous and safety-critical operations trates the general documentation flow. Some dif- for processing spacecraft systems and associated ferences exist depending on whether the payload is GSE. launching from the Eastern Range or the Western Test and Inspection Plans are required for the use Range. Contact the Delta Program Office for spe- of hoisting equipment and pressure vessels at the cific details. ranges. These plans describe testing methods, analy- Each Air Force and NASA safety agency has a ses, and maintenance procedures utilized to ensure requirement for submittal of documentation for compliance with 127-1 requirements. emitters of ionizing and nonionizing radiation. Re- Diligent and conscientious preparation of the quired submittals depend on the location, use, and required safety documentation cannot be overem- type of emitter and may consist of forms and/or phasized. Each of the USAF launch range support analyses specified in the pertinent regulations and organizations retains final approval authority over instructions. all hazardous operations that take place within its An RF ordnance hazard analysis must be per- jurisdiction. Therefore, the spacecraft agency should formed, documented, and submitted to confirm that consider the requirements of the 127-1 manual and the spacecraft systems and the local RF environ-

Table 9-1. Safety Document Applicability Safety Document Launch Payload EWR 127-1 KMI 1710.1G Astrotech SOP 1988 Astrotech SOP 1994 Site Type Reference A Reference B Reference C Reference D CCAS NASA X X Commercial X X VAFB NASA X X Commercial X X X

9-2 H30575.6 Spacecraft Agency

Distribution When NASA Facilities Are Used NASA GSFC/OLS

MDA/HB Review

ER NASA/KSC First SLS 45 SW/SESM MDA/CCAS Review/Approve Review Review/Approval Review

WR NASA/KSC-VAFB 30 SW/SES MDA/VAFB Review/Approve Review/Approval Review

Figure 9-1. General Safety Documentation Flow for Commercial Missions ment present no hazards to ordnance on the space- built, and tested to meet minimum factor of safety craft or launch vehicle. requirements (ratio between operating pressure and Each processing procedure that includes hazard- design burst pressure). All-metal vessels are acous operations must have a written procedure ap- ceptable, provided they are designed with a mini- proved by Space Wing safety (and NASA safety for mum calculated burst pressure of four times maxi- NASA facilities). Those that involve MDA person- mum allowable operating pressure, and that they nel or integrated operations with the launch vehicle also meet the requirements of EWR 127-1, Appen- must also be approved by MDA Test and Opera- dix 3C. MDA requires a minimum factor of safety tional Safety. of 2 to l for all pressure vessels which will be

9.3 HAZARDOUS SYSTEMS AND pressurized in the vicinity of MDA personnel. An OPERATIONS alternative approach exists for metal tanks designed, The requirements cited in the Space Wing safety built, and verified according to MIL-STD-1522A, regulations apply for hazardous systems and opera- Figure 2, Approach A, which are also acceptable for tions. However, MDA safety requirements are, in MDA personnel exposure. some cases, more stringent than those of the launch Increasing use of composite pressure vessels in range. The design and operations requirements gov- spacecraft propellant systems has led to the adop- erning activities involving MDA participation are tion of revised safety factor requirements for these discussed in the following paragraphs. types of pressure vessels. Safety factors applicable

9.3.1 Operations Involving Pressure to composite pressure vessels differ from those Vessels (Tanks) applicable to all-metal tanks. The required safety In order for MDA personnel to be exposed to factor varies according to the design and construc- pressurized vessels, the vessels must be designed, tion of the tanks. The applicable safety factor for 9-3 vessels of composite construction is dependent upon containing propellants must be capable of being the percentage of load carried by the overwrap. drained and then flushed and purged with inert MDA requirements are aligned with the launch fluids should a leak or other contingency require range requirements per EWR 127-1. Analyses and propellant offloading to reach a safe state. Space- test documentation verifying the pressure vessel craft designs should consider the number and place- safety factor must be included in the spacecraft ment of drain valves to maintain accessibility by safety documentation. technicians in Propellant Handler's Equipment (PHE Any operation that requires pressurization above suits) throughout processing. Coordinate with the levels that would maintain the required safety factor Delta Program Office to ensure that access can be must be conducted remotely (no personnel expo- accomplished while the payload fairing is in place sure) and requires a minimum 5-minute stabiliza- and that proper interfaces can be made with Delta tion period prior to personnel exposure. equipment and facilities.

9.3.2 Nonionizing Radiation 9.3.4 Safing of Ordnance The spacecraft nonionizing radiation systems are Manual ordnance safing devices (S&A or safing/ subject to the design criteria in the USAF and KSC arming plugs) for Range Category A ordnance are manuals and the special Delta-imposed criteria as also required to be accessible with the payload follows: fairing installed. Consideration should be given to ■ Systems producing nonionizing radiation will be placing such devices so that they can reached through designed and operated so that the hazards to person- fairing openings and armed as late in the count- nel are at the lowest practical level. down as possible and safed in the event of an ■ McDonnell Douglas employees are not to be aborted/scrubbed launch if required. Early coordi- exposed to nonionizing radiation above 1 mW/cm2 nation with the Delta Program Office is needed to averaged over any 1-minute interval. Safety docu- ensure that the required fairing access door(s) can mentation shall include the calculated distances at be provided.

which a level of 10 mW/cm2 (194 v/m) occurs (to 9.4 WAIVERS meet the USAF requirement) and the distances at Space Wing safety organizations discourage the which a level of 1 mW/cm2 (61 v/m) occurs (to meet use of waivers. They are normally granted only for the MDA requirement) for each emitter of nonion- spacecraft designs that have a history of proven izing radiation. This requirement is separate and safety. After a complete review of all safety re- distinct from the requirement to submit the radia- quirements, the spacecraft agency should deter- tion source documentation mentioned in Paragraph mine if waivers are necessary. A waiver or Meets 9.2. Intent Certification (MIC) request is required forany safety-related requirement that cannot be met. If a 9.3.3 Liquid Propellant Offloading non-compliant condition is suspected, coordinate Space Wing Safety Regulations require that liq- with the appropriate Space Wing Safety organiza- uid propellants aboard spacecraft be able to be tion to determine whether a Waiver or Meets Intent removed from tanks during any stage of prelaunch Certification will be required. Requests for waivers processing. Any tanks, piping, or other components shall be submitted prior to implementation of the 9-4 safety-related design or practice in question. Waiver Wing safety organizations determine when a waiver or MIC requests must be accompanied by sufficient or MIC is required and have final approval of all substantiating data to warrant consideration and requests. No guarantees can be made that approval approval. It should be noted that the USAF Space will be granted.

9-5 Appendix A Delta Missions Chronology Spacecraft Vehicle Delta Launch weight desig- Launch no. Mission date (kg/lb) Orbit parameters nation site *1 Echo I 05-13-60 62/137 1,667 x 1,667 km x 48 deg (900 x 900 nmi) Delta ETR 2 Echo IA 08-12-60 62/137 1,667x 1,667 km x 47 deg (900 x 900 nmi) Delta ETR 3 Tiros II 11-23-60 127/280 704 x 704 km x 48 deg (380 x 380 nmi) Delta ETR 4 Explorer X 03-25-61 35/78 174 x 222,240 km x 33 deg (94 x 120,000 nmi) Delta ETR 5 Tiros III 07-21-61 102/225 704 x 704 km x 48 deg (380 x 380 nmi) Delta ETR 6 Explorer XII 08-16-61 38/83 296 x 87,044 km x 33 deg (160 x 47,000 nmi) Delta ETR 7 Tiros IV 02-08-62 129/285 704 x 704 km x 48 deg (380 x 380 nmi) Delta ETR 8 OSO-A 03-07-62 208/458 556 x 556 km x 33 deg (300 x 300 nmi) Delta ETR 9 Ariel (UK) 04-26-62 62/136 370 x 1,019 km x 55 deg (200 x 550 nmi) Delta ETR 10 Tiros V 06-19-62 102/225 648 x 648 km x 58 deg (350 x 350 nmi) Delta ETR

11 Telstar I 07-10-62 77/170 926 x 5,556 km x 45 deg (500 x 3000 nmi) Delta ETR 12 Tiros V 09-18-62 129/285 648 x 648 km x 58 deg (350 x 350 nmi) Delta ETR 13 Explorer XIV 10-02-62 40/89 296 x 87,044 km x 33 deg (160 x 47,000 nmi) A ETR 14 Explorer XV 10-27-62 44/98 278 x 16,668 km x 18 deg (150 x 9,000 nmi) A ETR 15 Relay I 12-13-62 78/172 1296 x 7,408 km x 48 deg (700 x 4,000 nmi) B ETR 16 Syncom I 02-13-62 68/150 Synchronous transfer B ETR 17 Explorer XVI 04-02-63 186,409 250 x 898 km x 58 deg (135 x 485 nmi) B ETR 18 Telstar II 05-07-63 79/175 926 x 10,556 km x 43 deg (500 x 5,700 nmi) B ETR 19 Tiros VII 06-19-63 129/285 657 x 657 km x 58 deg (355 x 355 nmi) B ETR 20 Syncom II 07-26-63 67/147 Synchronous stationary B ETR

21 IMP-A 11-26-63 63/138 194 x 277,800 km x 33 deg (105 x 150,000 nmi) B ETR 22 Tiros VIII 12-21-63 129/285 685 x 685 km x 59 deg (370 x 370 nmi) B ETR 23 Relay II 01-21-64 83/184 2,130 x 7,408 km x 47 deg (1150 x 4,000 nmi) B ETR *24 Ionosphere Beacon 03-19-64 52/115 1,189 x 1,198 km x 71 deg (642 x 647 nmi) B ETR 25 Syncom III 08-19-64 66/145 Synchronous transfer D ETR 26 IMP-B 10-03-64 61/135 194 x 203,740 km x 33 deg (105 x 110,011 nmi) C ETR 27 Explorer XXVI 12-21-64 46/101 322 x 25,524 km x 20 deg (174 x 13,782 nmi) C ETR 28 Tiros IX 01-22-65 137/301 741 x 741 km x 98 deg (400 x 400 nmi) C ETR 29 OSO-B 02-03-65 249/548 556 x 556 km x 33 deg (300 x 300 nmi) C ETR 30 Early Bird 04-06-65 68/149 Synchronous transfer D ETR

31 IMP-C 05-29-65 58/128 189 x 222,240 km x 33 deg (102 x 120,000 nmi) C ETR 32 Tiros X 07-01-65 136/300 796 x 800 km x 99 deg (430 x 432 nmi) C ETR *33 OSO-C 08-25-65 281/619 556 x 556 km x 33 deg (300 x 300 nmi) C ETR 34 GEOS-A 11-06-65 174/384 1,111 x 1,483 km x 59 deg (600 x 801 nmi) E ETR 35 Pioneer-A 12-16-65 63.5/140 Heliocentric orbit, 0.984 x 0.829 AU E ETR 36 Tiros OT-3 02-03-66 129/285 737 x 756 km x 98 deg (398 x 408 nmi) C ETR 37 Tiros OT-2 02-28-66 148/326 1,324 x 1,396 km circular x 101 deg (715 x 743 nmi) E ETR 38 AE-B 05-25-66 223/492 270 x 1,204 km x 64 deg (146 x 650 nmi) C ETR 39 AIMP-D 07-01-66 93/206 6,667 x 555,296 km x 29 deg (3,600 x 299,836 nmi) E ETR 40 Pioneer-B 08-17-66 62.5/138 Heliocentric orbit 1.134 x 1.011 AU E ETR

41 TOS-A 10-02-66 148/326 1,400 x 1,420 km x 101 deg (756 x 767 nmi) E WTR 42 Intelsat II (F-1) 10-26-66 162/357 Synchronous transfer E ETR 43 BIOS-A 12-14-66 424.5/936 315 km circular x 33.5 deg (170 nmi) ETR 44 Intelsat II (F-2) 01-11-67 163/360 Synchronous transfer E ETR 45 TOS-B 01-26-67 129/285 1,432 x 1,439 km x 101 deg (773 x 777 nmi) E WTR 46 OSO-E1 03-08-67 283.5/625 556 km circular x 33 deg (300 nmi) C ETR 47 Intelsat II (F-3) 03-22-67 163/360 Synchronous transfer E ETR 48 TOS-C 04-20-67 148/326 1,433 x 1,439 km circular x 101 deg (774 x 777 nmi) E WTR 49 IMP-F 05-24-67 74/163 259 x 225,944 km x 66.5 deg (140 x 122,000 nmi) E WTR 50 AIMP-E 07-19-67 104/230 Lunar orbit E ETR

51 BIOS-B 09-07-67 433/955 315 km circular x 33.5 deg (170 nmi) G ETR 52 Intelsat II (F-4) 09-27-67 162/357 Synchronous transfer E ETR 53 OSO-D 10-18-67 290/640 556 km circular x 33 deg (300 nmi) C ETR

A-1 Appendix A Delta Missions Chronology (Continued) Spacecraft Vehicle Delta Launch weight desig- Launch no. Mission date (kg/lb) Orbit parameters nation site 54 TOS-D 11-10-67 148/326 1,463 km circular x 102 deg (790 nmi) E WTR 55 Pioneer-C (secondary 12-13-67 66/145 Heliocentric orbit, 1.100 x 0.987 AU E ETR payload; TTS and Cal-NCE) 56 GEOS-B (secondary 01-11-68 209.5/462 1,100 x 1,575 km x 106 deg (594 x 850 nmi) E WTR payload: Cal-NCE) 57 RAE-A 07-04-68 276/608 639 x 5,880 km x 121 deg (345 x 3,175 nmi) J WTR 58 TOS-E 08-16-68 157/347 1,464 km circular x 102 deg (790 nmi) N WTR *59 Intelsat III-A 09-18-68 291/641 Synchronous transfer M ETR

60 Pioneer-D (secondary 11-08-68 67/147 Heliocentric orbit, 0.991 x 0.755 AU E ETR payload: TTS) 24/53 61 HEOS-A 12-05-68 108/239 444 x 212,980 km x 28.3 deg (240 x 115,000 nmi) E ETR 62 TOS-F 12-15-68 154/340 1,463 km circular x 102 deg (790 nmi) N WTR 63 Intelsat III-C 12-18-68 293/647 Synchronous transfer M ETR 64 OSO-F 01-22-69 293/646 556 km circular x 33 deg (300 nmi) C ETR 65 ISIS-A 01-30-69 242/533 565 x 3,510 km x 88.5 deg (305 x 1,895 nmi) E WTR 66 Intelsat III-B 02-05-69 293/647 Synchronous transfer M ETR 67 TOS-G 02-26-69 156/345 1,463 km circular x 102 deg (790 nmi) N ETR 68 Intelsat III-D 05-21-69 293/647 Synchronous transfer M ETR 69 IMP-G 06-21-69 79/175 341 x 211,128 km x 83.8 deg (184 x 114,000 nmi) E WTR 70 BIOS-D 06-29-69 701/1546 370 km circular x 33.5 deg (200 nmi) N ETR

*71 Intelsat III-E 07-26-69 293/647 Synchronous transfer M ETR 72 OSO-G (secondary 07-26-69 293/647 Synchronous transfer M ETR payload: PAC) 08-09-69 289/638 556 km circular x 33 deg (300 nmi) N ETR 120/265 *73 Pioneer-E (secondary 08-27-69 67/148 Heliocentric orbit, 1.03 x 0.98 AU L ETR payload: TTS) 74 IDCSP/A 11-22-69 243/535 Synchronous transfer M ETR 75 Intelsat III-F 01-04-70 293/647 Synchronous transfer M ETR 76 Tiros-M (secondary 01-23-70 308/680 1,463 km circular 102 deg (790 nmi) N-6 WTR payload: Oscar) 25/55 77 NATO-A 03-20-70 243/535 Synchronous transfer M ETR 78 Intelsat III-G 04-22-70 293/647 Synchronous transfer M ETR 79 Intelsat III-H 07-23-70 293/647 Synchronous transfer M ETR 80 IDCSP/A-B 08-19-70 243/535 Synchronous transfer M ETR

81 ITOS-A 12-11-70 308/680 1,463 km circular x 102 deg (790 nmi) N-6 WTR 82 NATO-B 02-03-71 243/535 Synchronous transfer M ETR 83 IMP-I 03-13-71 288/635 237 x 212,928 km x 28.8 deg (128 x 114,972 nmi) M-6 ETR 84 ISIS-B 04-01-71 264/582 1,402 km x 88.7 deg (757 nmi) E WTR 85 OSO-H (secondary 09-29-71 636/1403 556 km circular x 33 deg (300 nmi) N ETR payload: TETR) *86 ITOS-B 10-21-71 309/682 1,463 km circular x 102 deg (790 nmi) N-6 WTR 87 HEOS-A2 01-31-72 117/258 409 x 245,149 km x 90 deg (221 x 132,370 nmi) L WTR 88 TD-1 03-11-72 476/1050 550 km sun-synchronous (297 nmi) N WTR 89 ERTS-A 07-23-72 939/2070 917 km sun-synchronous (495 nmi) 900 WTR 90 IMP-H 09-22-72 390/860 241 km x 38 Earth radii x 28.7 deg (130 nmi) 1604 ETR

91 ITOS-D (secondary 10-15-72 337/742 1,463 km sun-synchronous (790 nmi) 300 WTR payload: Oscar) 18.1/40 92 Telesat-A 11-10-72 562/1238 Synchronous transfer 1914 ETR 93 Nimbus-E 12-10-72 803/1770 1,111 km sun-synchronous (600 nmi) 900 WTR 94 Telesat-B 04-20-73 563/1242 Synchronous transfer 1914 ETR 95 RAE-B 06-10-73 333/735 Lunar orbit 1913 ETR *96 ITOS-E 07-16-73 340/749 1,463 sun-synchronous (790 nmi) 300 WTR 97 IMP-J 10-26-73 398/877 196 km x 37 Earth radii x 28.8 deg (106 nmi) 1604 ETR

A-2 Appendix A Delta Missions Chronology (Continued) Spacecraft Vehicle Delta Launch weight desig- Launch no. Mission date (kg/lb) Orbit parameters nation site 98 ITOS-F 11-06-73 340/749 1,520 km sun-synchronous (821 nmi) 300 WTR 99 AE-C 12-16-73 681/1500 157 x 4,315 km x 68 deg (85 x 2,330 nmi) 1900 WTR *100 Skynet IIA 01-18-74 435/960 Synchronous transfer 2313 ETR

101 Westar-A 04-13-74 574/1265 Synchronous transfer 2914 ETR 102 SMS-A 05-17-74 626/1379 Synchronous transfer 2914 ETR 103 Westar-B 10-10-74 574/1265 Synchronous transfer 2914 ETR 104 ITOS-G (secondary 11-15-74 340/749 1,463 km sun-synchronous (790 nmi) 2310 WTR payloads: Oscar 23/50 Intasat) 23/50 105 Skynet IIB 11-22-74 435/960 Synchronous transfer 2313 ETR 106 Symphonie-A 12-18-74 402/886 Synchronous transfer 2914 ETR 107 ERTS-B 01-22-75 964/2125 916 km sun-synchronous (495 nmi) 2910 WTR 108 SMS-B 02-06-75 630/1388 Synchronous transfer 2914 ETR 109 GEOS-C 04-09-75 337/742 848 km circular x 115 deg (458 nmi) 1410 WTR 110 Telesat-C 05-07-75 574/1265 Synchronous transfer 2914 ETR

111 Nimbus-F 06-12-75 907/2000 1,109 km sun-synchronous (599 nmi) 2910 WTR 112 OSO-1 06-21-75 1089/2400 556 km circular x 33 deg (300 nmi) 1910 ETR 113 COS B 08-08-75 277/611 348 x 99,993 km x 90 deg (188 x 53,992 nmi) 2913 WTR 114 Symphonie B 08-26-75 402/886 Synchronous transfer 2914 ETR 115 AE-D 10-06-75 675/1488 157 x 3,811 km x 90 deg (85 x 2,058 nmi) 2910 WTR 116 GOES-A 10-16-75 630/1388 Synchronous transfer 2914 ETR 117 AE-E 11-19-75 739/1630 159 x 2,998 km x 19 deg (85 x 1,619 nmi) 2910 ETR 118 RCA Satcom-A 12-12-75 868/1913 Synchronous transfer 3914 ETR 119 CTS 01-17-76 676/1490 Synchronous transfer 2914 ETR 120 Marisat-A 02-19-76 655/1445 Synchronous transfer 2914 ETR

121 RCA Satcom-B 03-26-76 868/1913 Synchronous transfer 3914 ETR 122 NATO III-A 04-22-76 700/1543 Synchronous transfer 2914 ETR 123 Lageos 05-04-76 408/900 5,900 km circular x 110 deg (3,186 nmi) 2913 WTR 124 Marisat-B 06-10-76 655/1445 Synchronous transfer 2914 ETR 125 Palapa-A 07-08-76 577/1273 Synchronous transfer 2914 ETR 126 ITOS-E-2 07-29-76 340/749 1,519 km sun-synchronous (820 nmi) 2310 WTR 127 Marisat-C 10-14-76 655/1445 Synchronous transfer 2914 ETR 128 NATO IIIB 01-28-77 700/1543 Synchronous transfer 2914 ETR 129 Palapa B 03-10-77 577/1273 Synchronous transfer 2914 ETR *130 ESRO-GEOS 04-20-77 574/1265 Synchronous transfer 2914 ETR

131 GOES-B 06-16-77 630/1388 Synchronous transfer 2914 ETR 132 GMS 07-14-77 669/1476 Synchronous transfer 2914 ETR 133 Sirio 08-25-77 398/877 Synchronous transfer 2313 ETR *134 OTS 09-13-77 865/1908 Synchronous transfer 3914 ETR 135 ISEE A/B 10-22-77 512/1128 280 x 140,320 km x 28.7 deg (151 x 75,767 nmi) 2914 ETR 136 Meteosat 11-22-77 663/1462 Synchronous transfer 2914 ETR 137 CS 12-14-77 675/1489 Synchronous transfer 2914 ETR 138 IUE 01-16-78 671/1479 167 x 46,344 km x 28.7 deg (90 x 25,024 nmi) 2914 ETR 139 Landsat C 03-05-78 940/2072 928 km (501 nmi) sun-synchronous 2910 WTR 140 BSE 04-07-78 676/1490 Synchronous transfer 2914 ETR

141 OTS-2 05-11-78 865/1908 Synchronous transfer 3914 ETR 142 GOES-C 06-16-78 635/1400 Synchronous transfer 2914 ETR 143 Esro-GEOS-2 07-14-78 574/1265 Synchronous transfer 2914 ETR 144 ISEE-C 08-12-78 478/1055 183 x 1,189,658 km x 28.78 deg (99 x 642,364 nmi) 2914 ETR 145 Nimbus G 10-24-78 938/2068 956 km sun-synchronous (516 nmi) 2910 WTR 146 NATO IIIC 11-18-78 700/1543 Synchronous transfer 2914 ETR

A-3 Appendix A Delta Missions Chronology (Continued) Spacecraft Vehicle Delta Launch weight desig- Launch no. Mission date (kg/lb) Orbit parameters nation site 147 Telesat-D 12-15-78 887/1956 Synchronous transfer 3914 ETR 148 Scatha 01-30-79 659/1452 185 x 43,226 km x 27.48 deg (100 x 23,340 nmi) 2914 ETR 149 Westar-C 08-09-79 574/1265 Synchronous transfer 2914 ETR 150 RCA-C 12-06-79 895/1974 Synchronous transfer 3914 ETR

151 SMM 02-14-80 2248/4956 574 km circular x 28.5 deg (310 nmi) 3910 ETR 152 GOES-D 09-09-80 833/1836 Synchronous transfer 3914 ETR 153 SBS-A 11-15-80 1083/2388 Synchronous transfer 3910 ETR 154 GOES-E 05-22-81 835/1841 Synchronous transfer 3914 ETR 155 DE-A/B 08-03-81 415/916 (A) 672 x 24,874 km x 90 deg (363 x 13,431 nmi) 3913 WTR (B) 305.6 x 1,300 km x 90 deg (165 x 702 nmi) 156 SBS-B 09-24-81 1085/2393 Synchronous transfer 3910 ETR 157 SME/Uosat 10-06-81 416/918 535 x 548.8 km x 97.5 deg (289 x 296 nmi) 2310 WTR 59.9/132 158 RCA-D 11-19-81 1082/2385 Synchronous transfer 3910 ETR 159 RCA-C 01-15-82 1082/2385 Synchronous transfer 3910 ETR

160 Westar IV 02-25-82 1108/2442 Synchronous transfer 3910 ETR

161 Insat-1A 04-10-82 1153/2541 Synchronous transfer 3910 ETR 162 Westar-V 06-08-82 1108/2442 Synchronous transfer 3910 ETR 163 Landsat-D 07-16-82 1972/4348 689 x 696 km x 98.3 deg (372 x 376 nmi) 3920 WTR 164 Telesat-F 08-26-82 1238/2730 Synchronous transfer 3920 ETR 165 RCA-E 10-27-82 1110/2447 Synchronous transfer 3924 ETR 166 IRAS/PIX II 01-25-83 1072/2375 906 x 909 km x 99.1 deg (489 x 491 nmi) 3910 WTR 167 RCA-F 04-11-83 1121/2472 Synchronous transfer 3924 ETR 168 GOES-F 04-28-83 835/1841 Synchronous transfer 3914 ETR 169 EXOSAT 05-26-83 510/1124 344 x 188,295 km x 72.5 deg (185.8 x 101,679 nmi) 3914 WTR 170 Galaxy-A 06-28-83 1218/2685 Synchronous transfer 3920 ETR

171 Telstar-3A 07-28-83 1218/2685 Synchronous transfer 3920 ETR 172 RCA-G 09-08-83 1121/2472 Synchronous transfer 3920 ETR 173 Galaxy-B 09-22-83 1218/2685 Synchronous transfer 3920 ETR 174 Landsat-D/Uosat 03-01-84 1973/4349 689 x 696.5 km x 98.3 deg (372 x 376 nmi) 3920 WTR 175 AMPTE 08-16-84 997/2198 939.4 x 113,417 km x 27 deg (507 x 61,245 nmi) 3924 ETR 176 Galaxy-C 09-21-84 1218/2685 Synchronous transfer 3920 ETR 177 NATO III-D 11-13-84 760/1675 Synchronous transfer 3914 ETR *178 GOES-G 05-03-86 838/1848 Synchronous transfer 3914 ETR 180 DM43 09-05-86 2495/5500 222 km circular x 28.6 deg (120 nmi) 3920/ ETR PAS

179 GOES-H 02-26-87 841/1853 Synchronous transfer 3924 ETR 182 Palapa-B2P 03-20-87 1244/2742 Synchronous transfer 3920 ETR 181 Thrusted Vector 02-08-88 1574/3470 222 x 333 km x 28.6 deg (120 x 180 nmi) 3910 ETR 184 Navstar II-1 GPS 02-14-89 1657/3645 20,183 km circular x 55.0 deg (10,898 nmi) 6925 ETR 183 Delta Star 03-24-89 2637/5800 270 nmi circular 3920 ETR 185 Navstar II-2 GPS 06-10-89 1657/3645 20,183 km circular x 55.0 deg (10,898 nmi) 6925 ETR 186 Navstar II-3 08-18-89 1664/3670 20,183 km circular x 55.0 deg incl (10,898 nmi) 6925 ETR 187 BSB R1 09-27-89 1233/2719 Synchronous transfer 4925 ETR 188 Navstar II-4 10-21-89 1664/3670 20,183 km/55.0 deg incl (10,898 nmi) 6925 ETR 189 COBE 11-18-89 2203/4857 900 km (486 nmi) circular 5920 WTR 190 Navstar II-5 12-11-89 1664/3670 20,183 km/55.0 deg incl (10,898 nmi) 6925 ETR

191 Navstar II-6 01-24-90 1664/3670 20,183 km/55.0 deg incl (10,898 nmi) 6925 ETR 192 Losat 02-14-90 2449/5400 546 km circular (295 nmi) 6920-8 ETR 193 Navstar II-7 03-25-90 1664/3670 20,183 km/55.0 deg incl (10,898 nmi) 6925 ETR 194 Palapa B2R 04-13-90 1241/2736 Synchronous transfer 6925-8 ETR 195 Rosat 06-01-90 2440/5380 580 km circular (313 nmi) 6920-10 ETR

A-4 Appendix A Delta Missions Chronology (Continued) Spacecraft Vehicle Delta Launch weight desig- Launch no. Mission date (kg/lb) Orbit parameters nation site 196 Insat 06-12-90 1293/2851 Synchronous transfer 5925 ETR 197 Navstar II-8 08-02-90 1664/3670 20,183 km circular x 55.0 deg (10,898 nmi) 6925 ETR 198 BSB-R2 08-17-90 1233/2719 Synchronous transfer 6925 ETR 199 Navstar II-9 10-01-90 1664/3670 20,183 km circular x 55.0 deg (10,898 nmi) 6925 ETR 200 Inmarsat 2-F1 10-30-90 1370/3020 Synchronous transfer 6925 ETR

201 Navstar II-10 11-26-90 1882/4150 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ETR 202 NATO IVA 01-07-91 1434/3161 Synchronous transfer 7925 ETR 203 Inmarsat 2-F2 03-08-91 1385/3054 Synchronous transfer 6925 ETR 204 ASC-2 04-12-91 1358/2994 Synchronous transfer 7925 ETR 205 Aurora II 05-29-91 1338/2950 Synchronous transfer 7925 ETR 206 Navstar II-11/LOSAT 07-03-91 1882/4150 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ETR 207 Navstar II-12 02-23-92 1882/4150 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ER 208 Navstar II-13 04-09-92 1882/4150 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ER 209 Palapa B4 05-13-92 1252/2761 Synchronous transfer 7925-8 ER 210 EUVE 06-07-92 3249/7165 528 km circular (285 nmi) 6920-10 ER

211 Navstar II-14 07-07-92 1882/4150 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ER 212 Geotail/DUVE 07-24-92 1009/2225 349,635 x 182 km x 28.7 deg (188,788 x 98.5 nmi) 6925 ER 213 Satcom C4 08-31-92 1402/3092 Synchronous transfer 7925 ER 214 Navstar II-15 09-09-92 1882/4150 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ER 215 Kopernikus 10-12-92 1411/3111 Synchronous transfer 7925 ER 216 Navstar II-16 11-21-92 1882/4150 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ER 217 Navstar II-17 12-18-92 1882/4150 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ER 218 Navstar II-18 02-02-93 1882/4150 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ER 219 GPS-1/SEDS-1 03-29-93 1882/4150 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ER 220 GPS 2 05-12-93 1882/4150 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ER

221 GPS-3/PMG 06-26-93 1882/4150 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ER 222 GPS-4 08-30-93 1882/4150 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ER 223 GPS-5 10-26-93 1882/4150 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ER 224 NATO-IVB 12-07-93 1434/3161 Synchronous transfer 7925 ER 225 Galaxy I-R 02-19-94 1400/3085 Synchronous transfer 7925-8 ER 226 GPS-6/SEDS-2 03-09-94 1882/4150 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ER 227 Wind 11-01-94 1250/2756 495,193 x 186.7 km x 28.75 deg (267,384 x 100.8 nmi) 7925 ER **228 Koreasat I 08-05-95 1447/3190 Synchronous transfer 7925 ER 229 Radarsat/SURFSAT 11-04-95 2865/6317 784 km x 800 km x 98.6 deg (423 x 432 nmi) 7920-10 WR 230 XTE 12-30-95 3030/6680 580 km circular x 23 deg (313.1 nmi) 7920-10 ER

231 Koreasat II 01-04-96 1457/3212 Synchronous transfer 7925 ER 232 NEAR 02-17-96 806/1778 Hyperbolic Escape (C=25.7) 7924-8 ER 233 POLAR 02-24-96 1301/2868 185 x 50,719 km x 86 deg (100 x 27,386 nmi) 7925-10 WR 234 GPS-7 03-28-96 1884/4154 20,183 km circular x 55.0 deg (10,898 nmi) 7925 ER 235 MSX 04-24-96 2800/6173 903 km circular x 99.23 deg (487.5 nmi) 7920-10 WR

*Vehicle Failure **One GEM solid motor carried to Stage I/II separation resulting in transfer orbit apogee outside of these sigma limits. Spacecraft fuel had to be expended to place satellite on station at the expense of lifetime. M029-App I T039-4/19/96 05:33 PM

A-5 (3) through or within 5 nautical miles Appendix B (horizontal or vertical) of the nearest edge NATURAL AND TRIGGERED LIGHTNING LAUNCH COMMIT CRITERIA of cumulus clouds with tops higher than the – 10.0 degree C level. (4) through or within 10 nautical miles The Delta launch vehicle will not be launched (horizontal or vertical) of the nearest edge if any of the following criteria are not met. Even of any cumulonimbus or thunderstorm when these constraints are not violated, if any cloud, including non-transparent parts of its other hazardous weather conditions exist, the anvil. Launch Weather Officer will report the threat to (5) through or within 10 nautical miles the Launch Director. The Launch Director may (horizontal or vertical) of the nearest edge hold at any time based on the instability of the of a non-transparent detached anvil for the weather. first hour after detachment from the parent A. Do not launch if any type of lightning is thunderstorm or cumulonimbus cloud. detected within 10 nautical miles of the planned Note:“Cumulus” does not include Altocumulus or flight path within 30 minutes prior to launch, unless Stratocumulus. the meteorological condition that produced the lightning has moved more than 10 nautical miles away C. Do not launch if, for ranges equipped with a from the planned flight path. surface electric field mill network, at any time B. Do not launch if the planned flight path will during the 15 minutes prior to launch time, the carry the vehicle: absolute value of any electric field intensity mea- (1) through a cumulus cloud with its top surement at the ground is greater than 1000 V/m between the + 5.0 degree C and - 5.0 degree within 5 nautical miles of the flight path unless: C level unless: (a) the cloud is not producing precipitation; (1) there are no clouds within 10 nautical miles -AND- of the flight path except: (b) the horizontal distance from the farthest (a) transparent clouds, edge of the cloud top to at least one -OR- working field mill is less than the (b) clouds with tops below the + 5.0 degree altitude of the – 5.0 degree C level, or C level that have not been associated 3 nautical miles, whichever is smaller; with convective clouds with tops above -AND- the – 10.0 degree C level within the (c) all field mill readings within 5 nautical past three hours, miles of the flight path are between – -AND- 100 V/m and + 1000 V/m for the (2) a known source of electric field (such as preceding 15 minutes. ground fog) that is occurring near the sensor, and that has been previously determined and (2) through cumulus clouds with tops higher documented to be benign, is clearly causing than the – 5.0 degree C level. the elevated readings. B-1 Note: For confirmed failure of the surface field of the flight path during the 15 minutes mill system, the countdown and launch may preceding the launch time; continue, since the other lightning launch commit -AND- criteria completely describe unsafe meteorological (3) The start of the 3-hour period is reckoned conditions. as follows: D. Do not launch if the flight path is through a (a) DETACHMENT. If the cloud detaches vertically continuous layer of clouds with an overall from the parent cloud: the 3-hour depth of 4,500 feet or greater where any part of the period begins at the time when cloud clouds is located between the 0.0 degree C and - detachment is observed or at the time 20.0 degree C levels. of the last detected lightning discharge E. Do not launch if the flight path is through any (if any) from the detached debris cloud, clouds that: whichever is later. (1) extend to altitudes at or above the 0.0 degree (b) DECAY or DETACHMENT UNCER- C level, TAIN. If it is not known whether the -AND- cloud is detached or the debris cloud (2) are associated with disturbed weatherthat is forms from the decay of the parent producing moderate (29 dBz) or greater cloud: the 3-hour period begins at the precipitation within 5 nautical miles of the time when the parent cloud top decays flight path. to below the altitude of the –10 degree F. Do not launch if the flight path will carry the C level, or at the time of the last vehicle: detected lightning discharge (if any) (1) through any non-transparent thunderstorm from the parent cloud or debris cloud, or cumulonimbus debris cloud during the whichever is later. first 3 hours after the debris cloud formed G. Good Sense Rule: Even when constraints are from the parent cloud. not violated, if hazardous conditions exist, the (2) within 5 nautical miles (horizontal or Launch Weather Officer will report the threat to the vertical) of the nearest edge of a non- Launch Director. The Launch Director may hold at transparent thunderstorm or cumulonimbus any time based on the weather threat. debris cloud during the first 3 hours after H. Definitions/Explanations the debris cloud formed from a parent cloud (1) Anvil: Stratiform or fibrous cloud produced unless: by the upper level outflow or blow-off from (a) there is at least one working field mill thunderstorms or convective clouds. within 5 nautical miles of the debris (2) Cloud Edge: The visible cloud edge is cloud; preferred. If this is not possible, then the -AND- 10 dBz radar cloud edge is acceptable. (b) all electric field intensity measurements (3) Cloud Layer: An array of clouds, not at the ground are between + 1000 V/m necessarily all of the same type, whose bases and – 1000 V/m within 5 nautical miles are approximately at the same level. Also,

B-2 multiple arrays of clouds at different The procedures used to assess the altitudes that are connected vertically by phenomena during launch countdowns shall cloud elements, e. g., turrets from one cloud be documented and implemented by the 45th to another. Convective clouds (e. g., clouds Weather Squadron. under Rule B above) are excluded from this (8) Electric Field (for surface-based electric definition unless they are imbedded with field mill measurements): The one-minute other cloud types. arithmetic average of the vertical electric (4) Cloud Top: The visible cloud top is field (Ez) at the ground, such as measured preferred. If this is not possible, then the by a ground-based field mill. The polarity 13 dBz radar cloud top is acceptable. of the electric field is the same as that of (5) Cumulonimbus Cloud: Any convective the potential gradient; that is, the polarity cloud with any part above the –20.0 degree of the field at the ground is the same as that C temperature level. of the charge overhead. (6) Debris Cloud: Any non-transparent cloud (9) Flight Path: The planned flight trajectory that has become detached from a parent including its uncertainties (‘error bounds’). cumulonimbus cloud or thunderstorm, or (10) Precipitating Cloud: Any cloud containing results from the decay of a parent cumulonimbus cloud or thunderstorm. precipitation, producing virga, or having (7) Documented: With respect to Rule C(2), radar reflectivity greater than 13 dBz. ‘documented’ means sufficient data has been (11)Thunderstorm: Any cloud that produces gathered on the benign phenomena to both lightning. understand it and to develop procedures to (12)Transparent: Synonymous with visually evaluate it, and the supporting data and transparent. Sky cover through which evaluation have been reported in a technical higher clouds, blue sky, stars, etc., may be report, journal article, or equivalent clearly observed from below. Also, sky publication. For launches at the Eastern cover through which terrain, buildings, etc., Range, copies of the documentation shall may be clearly observed from above. Sky be maintained by the 45th Weather cover through which forms are blurred, Squadron and KSC Weather Projects Office. indistinct, or obscured is not transparent.

B-3 Delta Mission Pueblo Checkout Production Facility Facility

Space Space Launch Launch Complex 17 Complex 2

Boeing Huntington Beach

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