f 

Copy No. SSD-CR-65-206 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 DOMED 

Contract AF04(695)-150 

TITAN III/MOL COMPATIBILITY STUDY (U) 

Performance Improvement Report 

Technical Summary 

September 1965 

Approved 

oeye 

L. J. dams, Technical Director 

i• , OLL(40,,,,,.r, 

(,. .) „,--,,, ., . i,., ,,,, ,,, ,,.-) 

1,1 

14 11 „'..4, ._ ) i:4 

,--)' 

L_VARTIN COMPANY 

Denver, Colorado 

Aerospace Division of Martin-Marietta, Corporation 

RECEIVED 

SEP 2 8 1965 

L M. WEEKS 
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SSD-CR-65-206 ii 

FOREWORD 

The data contained in this report summarize the results of a 

60-day study of performance improvements to the Titan III system. 

The performance improvements were achieved using larger solid 

rocket motors to replace the present 5-segment 120-in, diameter 

solid rocket motors of Stage O. These performance improvements 

were assessed against the requirements of the Manned Orbiting 

Laboratory (MOL) mission. Additional technical details for specific study areas are contained within individual tradeoff study 

reports summarized in the bibliography. 

This document is submitted under Item 1, Exhibit A, Task 5.13 

of Contract AF04(695)-150 in accordance with Line Item 3A-31 of 

Contractor Specification SSS-TIII-010 DRD (Rev 3), dated 15 April 

1963, and DSCN 1 thru 97. 
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SSD-CR-65-206 iii 

CONTENTS 

Page 

Foreword ii 

Contents . iii 

Summary . xvi 

thru 

xix 

I Introduction . I-1 

thru 

1-3 

II. Reference Configurations II-1 

A. Core Configurations . . . . , 11-1 

B. SRN Configurations ....... 11-27 

C, Payload Configurations . „ 11-32 

and 

11-33 

Performance and Trajectory Constraints III-1 

A. Performance Trade Studies . . III-1 

B, Performance Capability 111-7 

C. Vehicle Characteristics 111-52 

thru 

111-70 

IV, Control Systems IV-1 

A. Control System Stability IV-1 

B. TVC System Sizing IV-50 

C, Vibration Analysis IV-59 

V. Staging V-1 

A. Configurations Evaluated V-1 

B. Staging Criteria V-4 

C. Staging Sequence Considerations V-5 

D. Staging Sequence Acceleration Profile V-7 

E. Differential SRM Thrust Tailoff . V-11 

F. Stage 0/I Separation V-12 

thru 

V-22 
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VI. Solid Rocket Motors VI-1 

A. Reference Configurations VI-4 

B. SRM Constraints . Report Configuration VI-19 

C. SRM Contractor Data VI-30 

thru 

VI-56 

VII. Liquid Propulsion Systems .... VII-1 

A. Liquid Engine Data ....... VII-2 

B. Liquid Rocket Engine Changes VII-9 

C. Stage I/SRM Staging NPSH Study . • VII-l4 

D, Propellant Loading Inventory . . . VII-22 

thru 

VII-24 

VIII. Payload Length VIII-1 

A. Payload Description VIII-1 

D. Rigid Body Trajectory Results . . VIII-3 

C. Loads Analysis Assumptions VIII-6 

D. Maximum Airload and Length Analysis Results VIII-8 

E. Other Loads Analyses . . • • ..... VIII-17 

and 

VIII-18 

IX, Reliability and Crew Safety . . . • . . IX-1 

A. Ground Rules ........ . IX-1 

B. Approach IX-2 

C, Results ....... A A IX-4 

thru 

.IX-8 

X. Schedule ..... . • ...... • X-1 

A. Ground Rules X-1 

B. Titan III/MOL Booster Program Plan . X-2 

C. Equal Risk Schedule X-4 

D. Equal Risk Schedule for Stage 0 Development . . . ..... . . • X-11 

E, Summary X-17 
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XI, Bibliography . XI-1 

and 

X1-2 

Distribution 

Figure 

II-1 General Arranagement, MOL Compatibility 

Boosters 11-3 

11-2 Typical SRM Configuration 11-29 

11-3 SRM Acceleration vs Time 11-30 

11-4 SRM Thrust vs Time 11-31 

I1I-1 SRM Thrust Characteristics 111-4 

111-2 WTR Payload Capability 111-5 

111-3 Axial Acceleration Profile , ,,,,, 111-6 

111-4 Altitude-Velocity Plot, South WTR Launch 

with Transtage 111-9 

111-5 Altitude-Time Plot, South WTR Launch with 

Transtage III-10 

111-6 Dynamic Pressure-Time Plot, South WTR Launch 

with Transtage III-11 

111-7 Altitude-Velocity Plot, South WTR Launch 

without Transtage 111-12 

111-8 Altitude-Time Plot, South WTR Launch without 

Transtage 111-13 

111-9 Dynamic Pressure-Time Plot, South WTR Launch 

without Transtage 111-14 

III-10 Instantaneous Impact Points, 7 seg-120 with 

Transtage 111-15 

III-11 Instantaneous Impact Points, 2 CS-156 with 

Transtage 111-16 

111-12 Instantaneous Impact Points, 3 CS-156 with 

Transtage 111-17 

111-13 Instantaneous Impact Points, 7 seg-120 without Transtage 111-18 
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111-14 Instantaneous Impact Points, 2 CS-156 without Transtage 111-19 

111-15 Instantaneous Impact Points, 3 CS-156 without Transtage 111-20 

111-16 Reentry Abort Ceiling, South WTR Launch with 

Transtage 111-22 

111-17 Launch Trajectory Temperature Constraints, 

South WTR Launch with Transtage III-23 

111-18 Allowable gap vs Mach, Ejection Seat Escape 

System, South WTR Launch with Transtage . , 111-24 

111-19 Park Orbit Payload Capability, 7 seg-120 . . 111-28 

III-20 Payload Altitude Profile, 7 seg-120, 100-nmi Park Orbit 111-29 

111-21 Transtage Propellant Usage, 7 seg-120, 100- 

n-mi Park to Circular Orbits 111-30 

111-22 Transtage Propellant Usage, 7 seg-120, 100- 

n-mi Park to Elliptical Orbits 111-31 

111-23 Payload Altitude Profile, 7 seg-120, 100- 

n-mi Perigee Overspeed 111-32 

111-24 Transtage Propellant Usage, 7 seg-120, 100- 

n-mi Perigee to Circular Orbits 111-33 

111-25 Transtage Propellant Usage, 7-seg-120, 100- 

n-mi Perigee to Elliptical Orbits . • . . 111-34 

111-26 Park Orbit Payload Capability, 2 CS-156 

• " 

111-35 

111-27 Payload Altitude Profile, 2 CS-l56, 100-n-mi 

Park Orbit 111-36 

111-28 Transtage Propellant Usage, 2 CS-156, 100- 

n-mi Park to Circular Orbits ^ ^.... ^ 111-37 

111-29 Transtage Propellant Usage, 2 CS-156, 100- 

n-mi Park to Elliptical Orbits ..... 111-38 

111-30 Payload Altitude Profile, 2 CS-l56, 100-n-mi 

Perigee Overspeed 111-39 

111-31 Transtage Propellant Usage, 2 CS-156, 100- 

n-mi Perigee Overspeed to Circular Orbits 111-40 

111-32 Transtage Propellant Usage, 2 CS-156, 100- 

n-mi Perigee Overspeed to Elliptical 

Orbits 111-41 
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111-33 Park Orbit Payload Capability, 3 CS-156 . 111-42 

111-34 Payload Altitude Profile, 3 CS-156, 100-n-mi 

Park Orbit 111-43 

111-35 Transtage Propellant Usage, 3 CS-156, 100- 

n-mi Park to Circular Orbits 111-44 

111-36 Transtage Propellant Usage, 3 CS-156, 100- 

n-mi Park to Elliptical Orbits 111-45 

111-37 Payload Altitude Profile, 3 CS-156, 100-n-mi 

Perigee Overspeed 111-46 

111-38 Transtage Propellant Usage, 3 CS-156, 100- 

n-mi Perigee Overspeed to Circular Orbits 111-47 

111-39 Transtage Propellant Usage, 3 CS-156, 100- 

n-mi Perigee Overspeed to Elliptical 

Orbits 111-48 

111-40 Trajectory Shaping, Payload Increase, 

3 CS-l56 111-49 

111-41 Trajectory Shaping, Aerodynamic Heating, 

3 CS-156 111-50 

111-42 Trajectory Shaping, Maximum Dynamic Pressure, 3 CS-156 111-51 

111-43 Thrust-Time History, 7 seg-l20 111-53 

111-44 Thrust-Time History, 2 CS-156 111-54 

111-45 Thrust-Time History, 3 CS-156 111-55 

111-46 Axial Force Coefficient, 7 seg-120, 

Stage 0 111-67 

111-47 Axial Force Coefficient, 2 CS-156, Stage 0 111-68 

111-48 Axial Force Coefficient, 3 CS-156, 

Stage 0 111-69 

111-49 Axial Force Coefficient, Stage 1, All Vehicles 111-70 

IV-1 Yaw Structural Mode Frequencies; Flight 

Time = Liftoff, with Trarstage IV-3 

IV-2 Load Reduction by the PCS, with Transtage IV-4 

IV-3 Pitch Autopilot 1V-6 

IV-4 Typical First and Second Mode Shapes , • IV-8 
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IV-5 Frequency Response Plot for Stage 0; Pitch, 

T = 60 sec; 7 seg-120; 54.5-ft, 28,000-lb 

Payload ....... . . ^ . . ^ 1V-12 

IV-6 Frequency Response Plot for Stage 0; Yaw, 

T = 60 sec; 7 seg-120; 54.5-ft, 28,000-lb 

Payload ....... . . . . IV-13 

IV-7 Frequency Response Plot for Stage 0; Pitch, 

T = 60 sec; 7 seg-120; 75.0-ft, 28,000-lb 

Payload " " " * .. ' . 

IV-14 

IV-8 Frequency Response Plot for Stage 0; Yaw, 

T = 60 sec; 7 seg-120; 75.0-ft, 28,000-lb 

Payload IV-15 

1V-9 Frequency Response Plot for Stage 0; Pitch, 

T = 60 sec; 2 CS-156; 50,0-ft, 35,000-lb 

Payload 1V-16 

1V-10 Frequency Response Plot for Stage 0; Yaw, 

T = 60 sec; 2 CS-156; 50.0-ft, 33,000-lb 

Payload 1V-17 

1V-11 Frequency Response Plot for Stage 0; Pitch, 

T = 60 sec; 2 CS-l56; 65.0-ft, 33,000-lb 

Payload 1V-18 

1V-12 Frequency Response Plot for Stage 0; Yaw, 

T = 60 sec; 2 CS-156; 65.0-ft, 33,000-lb 

Payload 1V-19 

IV-l3 Frequency Response Plot for Stage 0; Pitch, 

T = 60 sec; 3 CS-156; 50.0-ft, 42,000-lb 

Payload 1V-20 

1V-14 Frequency Response Plot for Stage 0; Yaw, 

T = 60 sec; 3 CS-156; 50.0-ft, 42,000-lb 

Payload 1V-21 

1V-15 Frequency Response Plot for Stage 0; Pitch, 

T -= 60 sec; 3 CS-156; 72,0-ft, 42,000-16 

Payload IV-22 

IV-16 Frequency Response Plot for Stage 0; Yaw, 

T = 60 sec; 3 CS-156; 72.0-ft, 42,000-lb 

Payload IV-23 

IV-17 Frequency Response Plot for Stage 0; Pitch, 

T = 80 sec; 65.0-ft, 33,000-lb Payload; 

Sensors at Stations 235 and 887 r7-24 
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SSD-CR-65-206 ix 

IV-18 Frequency Response Plot for Stage 0; Yaw, 

T = 80 sec; 65,0-ft, 33,000-lb Payload; 

Sensors at Stations 235 and 887 •••—• IV-25 

IV-19 Frequency Response Plot for Stage 0; Pitch, 

T = 80 sec; 65.0-ft, 33,000-lb Payload; 

Sensors at Stations 112 and 887 1V-26 

IV-20 Frequency Response Plot for Stage 0; Yaw, 

T = 80 sec; 65.0-ft, 33,000-lb Payload; 

Sensors at Stations 112 and 887 IV-27 

1V-21 Frequency Response Plot for Stage I; Pitch/ 

Yaw, T = Start; 28,000-1b, 75-ft Payload „ IV-29 

1V-22 Frequency Response Plot for Stage 1; Pitch/ 

Yaw, I = Midnight (Start + 119 sec), Before 

Gain Change; 28,000-1b, 75-ft Payload . . IV-30 

IV-23 Frequency Response Plot for Stage I; Pitch/ 

Yaw, = Midflight (Start + 119 sec), After 

Gain Change; 28,000-1b, 75-ft Payload . . . IV-31 

IV-24 Frequency Response Plot for Stage I; Pitch/ 

Yaw, T = Burnout; 28,000-lb, 75-ft Payload . IV-32 

IV-25 Frequency Response Plot for Stage II; Pitch/ 

Yaw, I = Start; 7 seg-120; 65.0-ft, 28,000- 

lb Payload . • ..... .. • • • 1V-34 

IV-26 Frequency Response Plot for Stage II; Pitch/ 

Yaw, T = Midflight; 7 seg-120; 65.0-ft, 

28,000-lb Payload IV-35 

IV-27 Frequency Response Plot for Stage II; Pitch/ 

Yaw, I = Burnout; 7 seg-120; 65,0-ft, 

28,000-lb Payload IV-36 

1V-28 TVC Fluid Utilization, 7 seg-120 1V-52 

IV-29 TVC Fluid Utilization, 2 CS-l56 IV-53 

IV-30 TVC Fluid Utilization, 3 CS-156 IV-54 

V-1 Staging Sequence . . . V-6 

V-2 SRM Tailoff V-9 

V-3 Stage I Ultimate Longeron Load (Vehicle Station 1224) at Stage I Engine Start . . V-10 

V-4 Control Authority during SRN Tailoff (Minimum/Maximum SRN Tailoff, Late Core Start). ^ V-11 

V-5 7 seg-120 SRN Staging V-13 

V-6 2 CS-156 SRM Staging . . ..... V-16 
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V-7 3 CS-156 SRM Staging . . . . . . ...... V-19 

VI-1 Stage 0 Design (Loads Reference Configuration) . . . . . ............ . . VI-5 

V172 7 seg-120 SRM Reference Configuration . . . VI-7 

VI-3 Thrust vs Time, Reference Configurations . . VI-10 

VI-4 7 seg-120 Grain Design VI-11 

VI-5 2 CS-156 SRM Reference Configuration . . , VI-13 

VI-6 2 CS-156 SRM Grain Design VI-14 

V1-7 3 CS-156 SRM Reference Configuration . . • • VI-16 

VI-8 3 CS-156 SRN Grain Design V1-17 

VI-9 2 CS-156 SRM Constraints Document Configuration VI-21 

VI-10 3 CS-156 SRM Constraints Document Configuration VI-23 

VI-11 2 CS-156 SRM Constraints Document Configuration, Thrust vs Time . , , . . . , , . . . . VI-25 

VI-12 3 CS-156 SRM Constraints Document ConfiguraVI-18 Ignition and Tailoff Parameter Definitions , V1-37 

VI-19 Tailoff Vacuum Thrust Curves, 7 seg-120 . VI-38 

VI-20 Thrust Tailoff, LPC 2 CS-156 VI-39 

VI-21 Thrust Tailoff, LPC 3 CS-156 VI-39 

VI-22 Thrust Tailoff, TCC 2 CS-156 VI-39 

VI-23 Maximum Thrust Differential vs Tailoff 

Time V1-40 

VI-24 Relationship of Igniter Discharge to Ignition Delay . . . .......... . . • . V1-41 

VI-25 Nominal Ignition Transient, Large SRMs . . V1-41 

tion, Thrust vs Time VI-25 

VI-13 LPC 2 CS-156 Motor Performance (60°F) 

VI-14 ICC 2 CS-156 Motor Performance (60°F) . . VI-32 

VI-15 AGC 2 CS-I56 Motor Performance (60°F) 

VI-16 LPC 3 CS-156 Motor Performance (60°F) • . • VI-33 

VI-17 UTC 3 CS-156 Motor Performance (60°F) . . • VT 33 
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VI-26 TVG in UCC 7 seg-120, Side Force vs 

Voltage „ .. , ... ...... , VI-45 

VI-27 TVC in UTC 7 seg-120, Flow late vs Voltage V1-45 

VI-28 TVC in LPC 2 GS-156, Side Force. vs Voltage VI-46 

VI-29 TVC in LPC 2 CS-156, Flow Rate vs Voltage . VI-46 

VI-30 TVC in LPG 3 CS-156, Side Force vs Voltage . VI-47 

VI-31 TVC in LPC 3 CS-156, Flow Rate vs Voltage . VI-47 

VI-32 Maximum Thrust Vector Deflection vs Number 

of Injectant Valves VI-48 

V1-33 Typical Proposed Insulation Joints, 156-in, 

Diameter VI-53 

VI-34 SRM Net Forward Thrust Along Nozzle Centerline to Maintain Total Thrust Along Vehicle 

Centerline of 100,000 lb VI-54 

VI-35 Thrust Termination Evaluation, UTG 7 seg120 VI-54 

VI-36 Thrust Termination Evaluation, LPC 2 

CS-156 ..... ...... . . VI-55 

VI-37 Thrust Termination Evaluation, LPC 3 

CS-156 VI-55 

VII-1 Titan III/MOL Exhaust Plume at 220,000 ft VII-4 

VII-2 Preliminary Stage I Titan III/MOL Engine 

Assembly VII-7 

VII-3 Titan III/MOL Stage I Engine Exit Plane/ 

Vehicle Outside Skin Line Layout (15:1 Expansion Ratio) .... . ..... . . . VII-8 

VII-4 15:1 Stage I Engine Ablative Skirt VII-11 

V11-5 Proposed Improved Fuel Pump V1I-13 

VII-6 Titan III/MOL Stage I Oxidizer Minimum NPSH 

during SRM Staging (7 seg-120; 75°F Propellant Temperature) VII-18 

VII-7 Titan III/MOL Stage I Oxidizer Minimum NFS11. 

during SRM Staging (7 seg-120; 90°F Propellant Temperature) VII-l8 

VII-8 Titan III/MOL Stage I Oxidizer Minimum NPSH 

during SRM Staging (2 CS-156; 75°F Propellant Temperature) VII-19 
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VII-9 Titan III/MOL Stage I Oxidizer Minimum NPSH 

during SRN Staging (2 CS-156; 90°F Propellant Temperature) VII-19 

VII-10 Titan III/MOL Stage I Oxidizer Minimum NPSH 

during SRN Staging (3 CS-156; 75°F Propellant Temperature) VII-20 

VII-11 Titan III/MOL Stage I Oxidizer Minimum NPSH 

during SRN Staging (3 CS-156; 90°F Propellant Temperature) VII-20 

VII-12 Titan III/MOL Stage I Oxidizer Minimum NPSH 

during Staging (3 CS-156; 90°F Propellant. 

Temperature; Increased Lockup Pressure) . VII-21 

VIII-1 Airload Indicator vs Percentage of Maximum 

Wind VIII-5 

VIII-2 Launch Probability vs Design Wind VIII-7 

VIII-3 Ultimate Core Loads, 7 seg-120 . . • VIII-9 

VIII-4 Ultimate Core Loads, 2 CS-156 VIII-10 

VIII-5 Ultimate Core Loads, 3 CS-156 • • • • • VIII-11 

VIII-6 Allowable Payload Lengths, 99% Launch Probability VIII-12 

VIII-7 Allowable Payload Length vs Launch Probability . . ...... VIII-14 

X-1 Titan III/MOL Booster Program Plan X-3 

X-2 7 seg-120 with Transtage; 15:1 Stage I 

Engine Configuration Schedule X-5 

X-3 7 seg-120 without Transtage; 15:1 Stage I, 

Engine Configuration Schedule X-7 

X-4 2 CS-156 with Transtage; 8:1 Stage I Engine 

Configuration Schedule X-8 

X-5. 3 CS-156 with Transtage; 8:1 Stage I Engine 

Configuration Schedule X-10 

X-6 7 seg-120 SAM Development Schedule X-12 

X-7 2 CS-156 SRM Development Schedule X-14 

X-8 3 CS-I56 88M Development Schedule X-16 
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Table 

11-1 Reference Configurations II-1 

11-2 Summary of Configuration Modifications to 

the Core due to Performance Improvement 

Changes 11-5 

11-3 Core Structural Modification due to Performance Improvement Changes 11-6 

11-4 Configuration Summary of Crew Safety/Mission 

Success Modifications 11-17 

11-5 Core Structural Modifications due to Crew 

Safety/Mission Success Requirements . . . , 11-17 

11-6 Weight Summary 11-22 

11-7 Comparison Chart of Compartmental Vibration 

(grms) Levels 11-26 

11-8 SRM Characteristics 11-28 

11-9 Spacecraft Configuration 11-33 

III-1 Solid Rocket Motor Characteristics . , . ^ 111-3 

111-2 WTR Performance Summary, WTR Launch, Polar 

80-n-mi Circular Orbit 111-8 

111-3 Stage I Class Nominal, Expansion Ratio, 

15:1; Used with 7 seg-120 111-56 

111-4 Stage I Class Nominal, Expansion Ratio, 8:1; 

Used with 2 CS-156 and 3 CS-l56 111-57 

111-5 Stage II Class Nominal, Expansion Ratio, 

49.2:1; Used with All Configurations . . . 111-58 

111-6 Liquid Propellant Inventory 111-59 

111-7 Step 0 Performance Weight Summary with and 

without Transtage 111-61 

111-8 Step I Performance Weight Summary with 

Transtage 111-62 
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111-9 Step 1 Performance Weight Summary without 

Transtage „ . . ........... . 111-63 

1II-10 Step 2 Performance Weight Summary with 

Transtage 111-64 

III-11 Step 2 Performance Weight Summary without 

Transtage 111-65 

111-12 Step 3 Performance Weight Summary . . . . 111-66 

1V-1 Stage 0 Flight Control System Configurations . . . ......... . . . . . 

IV-2 Stage I Flight Control System Configurations IV-47 

IV-3 Stage II Flight Control System Configurations IV-48 

IV-4 Summary of Fluid Requirements IV-50 

IV-5 Injectant Requirements, 7 seg-120 IV-55 

IV-6 Injectant Requirements, 2 CS-156 IV-56 

IV-7 Injectant Requirements, 3 CS-156 IV-57 

V-1 Configurations Evaluated V-2 

V-2 Staging Mass Property Data . . . ... V-3 

V-3 Staging Sequence and Accelerations V-7 

V-4 Summary of Separation Plots V-12 

VI-1 SRM Reference Configuration Characteristics VI-8 

VI-2 SRM Performance Characteristics VI-9 

VI-3 Recommended 156-in. SRM Physical Characteristics VI-20 

VI-4 Recommended 156-in. SRM Performance Characteristics VI-20 

VI-5 SRM Internal Ballistics VI-34 

VI-6 SRM Performance Tolerances (60°F) VI-36 

VI-7 Tailoff Parameters VI-36 

VI-8 Ignition Parameters VI-42 

VI-9 TVC System Requirements VI-44 

VI-10 Nozzle Characteristics VI-49 

IV-37 
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VI-11 Case and Attachments Design Criteria , VI-50 

VI-12 Proposed Insulation Systems VI-52 

VI-13 Staging Motors Required VI-56 

VII-1 Stage I Engine Characteristics VII-3 

VII-2 Stage II Engine Characteristics VII-5 

VII-3 Stage III Engine Characteristics VII-5 

VII-4 Liquid Propellant Inventory, Titan III/MOL 

(Transtage Off-Loaded) VII-23 

VII-5 Liquid Propellant Inventory, Titan III/MOL 

(Transtage Fully Loaded) . . VII-24 

VIII-1 Nominal Payload Lengths Used for Loads Analyses (with-Transtage Configuration) . . . . VIII-2 

VIII-2 Trajectory Characteristics for the Critical 

Flight Conditions VIII-4 

VIII-3 Delta Structural Weight (lb) Associated with 

Core Redesign to Permit 99% Launch Probability VIII-15 

VIII-4 Summary of Possible Payload Length Increases 

(ft) VIII-16 

VIII-5 Gemini/Laboratory Module Bending Moments and 

Axial Loads for Critical Flight Conditions, 

Ultimate Loads VIII-18 

IX-1 Crew SafetytMission Aborts Summary IX-5 

IX-2 Mission Aborts by Flight Phase IX-6 

IX-3 Crew Safety/Mission Abort Summary by Subsystem Grouping IX-7 

IX-4 Effects of Step 0 Burning Time on Reliability IX-8 
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SUMMARY 

The purpose of the performance improvement study was to investigate the performance improvements that could be achieved on 

the current Titan III system by incorporating larger solid rocket 

motors (SRMs) to replace the present 5-segment 120-in, diameter 

motors for the Manned Orbiting Laboratory (MOL) mission. Three 

SRM options were considered for MOL missions launched from the 

Western Test Range (WTR) into low altitude polar orbits. Each 

option was considered both with and without use of the transtage 

through orbit injection. The SRM options were: 

1) 7 segment, 120-in. (7 seg-120); 

2) 2 center segment, 156-in. (2 CS-156). 

3) 3 center segment, 156-in. (3 CS-156). 

The 7 seg-120 configuration considered a 15:1 Stage I engine expansion ratio. The Stage I engines for the 156-in. SRM configurations, except as used for a staging analysis, considered only an 

8:1 expansion ratio engine. 

Study results indicate that minimum modification to the Titan 

III core stages is required for all SRMs. The aft longeron in 

Stage I must be strengthened to accept the 156-in. SRMs. The 

local vibration environment in the region of attachment for the 

longer SRMs (e.g., 7 seg-120 and 3 CS-156) will necessitate requalification, relocation, and redesign of some Titan III core 

components. Another forward staging rocket will be required for 

staging of the 156-in. SRMs. Net pump suction head (NPSH) requirements of the oxidizer feedline to the Stage I engine assembly can 

be met by increasing the tank lockup pressure without modifying 

the core for the 7 seg-120 SRM using the 15:1 expansion ratio 

engine. A relaxation in NPSH requirements to 35-ft head is required for the 156-in. SRMs using either the 8:1 or 15:1 expansion 

ratio engine. 

All SRM designs met the preselected vehicle compatibility constraints for liftoff acceleration (>1,6 g), maximum dynamic pressure 

(<900 psf), limit in-flight acceleration (<3,2g), and maximum 

aerodynamic heating (<100 x 106 ft-lb//ft2

). The SRM designs were 

based on state-of-the-art technology using conventional class 2 

propellants and liquid inject thrust vector control (TVC). Analysis of the injectant requirements showed that the Titan III tank 
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size is adequate to provide total injectant for worst-case conditions. SRM weights (per pair) varied from 1,364,000 (7 seg-120) 

to 2,537,000 lb for the 3 CS-156 option. For all configurations, 

total SRM vacuum impulse varied from 308,000,000 to 588,000,000 

lb-sec compared to the present 5 seg-120 SRM 210,000,000 lb-sec 

impulse. 

Payload performance into low altitude orbit was found to be 

almost directly proportional to Stage 0 vacuum impulse and to be 

increased by approximately 1000 lb when the transtage was not 

used. Guaranteed payload was 28,306, 35,030, and 42,109 lb for 

the 7 seg-120, 2 CS-156, and 3 CS-156 options, respectively, for 

the without-transtage configurations launched into an 80-n-mi polar 

orbit. East launch of these configurations into a 100-n-mi circular orbit from the Eastern Test Range (ETR) would give payload 

capability of 33,000, 41,000, and 49,500 lb, respectively. The 

minimum performance capability of the current Titan IIIC using 

the 15:1 Stage I engine and anticipated MOL equipment weights 

would be 22,206 lb into an 80-n-mi circular polar orbit from WTR 

without transtage. 

Analyses of the flight control system for each configuration 

option, across the range of specified maximum payload lengths, 

demonstrated that the current Titan III autopilot could be used. 

Modifications to increase the capacity of the adapter programer 

and flight control computer will be necessary to provide 207. additional capacity for any of the SRM options when used with the 

transtage. This increased capacity is not required for the withouttranstage configurations since adequate capacity for the additional rate gains necessitated by the longer burn time SRM options 

can be obtained when the orbital coast requirement is deleted. A 

new snubber will be required for Stage I for the 7 seg-120 configuration option with the 15:1 Stage I engine. Adequate stability margins were demonstrated for all configurations. Relocation of the inertial measurement unit (EMU), forward rate gyro, 

and lateral acceleration sensor will be required for all configurations. The load-relief accelerometer system was used for all 

flight simulations with effectiveness ranging between 28 and 35% 

depending on the specific configurations. Yaw structural mode 

frequencies were not sensitive to payload length. 

Results of the rigid body loads analysis showed the maximum 

airload indicators (TV.'

max

) to be about 5000 lb-deg/ft

2 for all 

configurations regardless of payload length, and to be approximately the magnitude currently experienced on the Titan IIIC. 

UNCLASSIFIED 
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Worst wind azimuths occurred at either 245 or 272 deg azimuth for 

wind shear peak altitudes between 29,000 and 32,000 ft at about Mach 

1.40. It was determined that all configurations would carry MOL 

payload lengths of at least 51 ft without structural beefup of the 

core and in the worst wind condition. The 7 seg-120 configuration 

could transport MOL payload lengths of 57.8 and 74.7 ft for the 

with- and without-transtage options, respectively. Comparable 

lengths for the 2 CS-156 options were 51.3 ft with transtage and 

67.8 ft without transtage and, for the 3 CS-156 options, were 

53.0 ft with transtage and 71.2 ft without transtage. The analysis demonstrated that a 70-ft payload length could be carried with 

a >97% launch probability for all configurations without transtage 

and >65% for all configurations with transtage. Analyses of the 

required structural beefup to the core to accept longer payloads 

without wind placarding and restriction of launch probability 

showed that relatively small increases in core structural weights 

(,<i930 lb) would permit delivery/transport of all specified maximum overall length payloads; e.g., 70 ft with transtage and 82 ft 

without transtage. 

A preliminary reliability and crew safety study was performed 

to assess the impact of the performance and preliminary crew 

safety modifications on mission reliability and crew safety. 

Crew safety analysis is defined as a warning-time analysis only. 

Analysis to date indicates that all of the with-transtage configurations (130-n-mi orbital mission) meet the mission success 

goal of 0.94 and that the best without-transtage configuration 

prediction is 0.964 compared to a goal of 0.970 for an 80-n-mi 

orbital mission. Differences in the six configurations, however, 

are insignificant. Preliminary crew safety analysis indicates 

that none of the configurations meet the goal of less than 1800 

mission aborts per million flights with a warning time less than 

3 sec, The predictions indicate a range of 2202 to 2598 mission 

aborts with a warning time less than 3 sec. The SRNs are the 

largest contributor to this estimate with a range of 1213 to 

1364. 

During the performance improvement study, an equal-risk schedule was prepared for the six configurations. This review indicated 

that the 7 seg-120 configuration has the least risk schedule. The 

controlling item for the 7 Seg-120 configuration was the 15:1 Stage 

I liquid rocket engine (LRE) development, which would require outof-position installation of the production engine to meet the May 

1968 launch schedule. The 2 CS-156 and 3 CS-156 schedules indicated that a long-lead go-ahead for the case procurement and nozzle 

design would be required to meet the May 1968 launch schedule. It 

was concluded from the schedule review that all configurations, 

with the appropriate long-lead go-ahead, could meet the May 1968 

launch schedule with equal risk. 

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xix 

It was determined from this 60-day study that each of the configuration options resulted in a feasible system that could be developed against MOL schedule requirements with low technical risk. 

All proposed systems were based on present state-of-the-art technology. 

In view of the minimal diffetence in effect on the core for 

the various performance options, we have concluded that no basis 

exists for a Martin recommendation as to ultimate configuration 

selection. 
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I. INTRODUCTION 

In 1964, Martin Company, Denver Division, participated in preliminary program definition studies under contract to the Space 

Systems Division (SSD), Air Force Systems Command, to identify 

the system elements to comprise the Manned Orbiting Laboratory 

(MOL) system. The compatibility of the Titan IIIC space launch 

system was assessed against the MOL mission requirements defined 

at that time. The compatibility assessment included studies of 

the performance capability of the Titan III with the present solid rocket motors (SRMs) and with larger SRMs. Payload performance was determined for low altitude orbits when launched from 

the Eastern Test Range (ETR) in an easterly direction and from 

the Western Test Range (WTR) into polar orbits. The use of the 

transtage as an integral part of the on-orbit portion of the MOL 

system was studied to establish changes required to the transtage 

for on-orbit times up to 30 days. The maximum length of payloads 

that could be transported on the current Titan IIIC were established for a family of cylindrical laboratory modules including 

a modified Gemini capsule. The requirements for new equipment 

were established as necessary to achieve increased mission success 

and improved crew safety using the Gemini capsule with and with, 

out a new launch escape tower. These studies were preliminary in 

nature, and the major emphasis was on the use of the current Titan 

IIIC system with 5-segment 120-in, diameter SRMs from ETR. 

Based on results of these preliminary program definition studies, 

SSD further defined the requirements for the MOL system, particularly as they affected the choice of the Titan III system configuration to be used for the mission. It was established that development of the full operational capability for the MOL system 

would require launch into polar orbits from the WTR with payload 

weights in excess of the current Titan IIIC capability. Laboratory module diameter was set at 126 in. for lengths up to 82 ft, 

The transtage was not to be used as a part of the on-orbit system 

and might be used as part of the system to orbit injection. Increased mission success was defined with associated requirements 

for modifications to core subsystems, The decision was made not 

to provide a forward mounted launch escape tower on the Gemini 

capsule, Additional necessary subsystems to provide adequate margins for crew safety were to be provided in the Titan III core 

stages through increased redundancy, improved reliability of components, and use in alternative modes of selected systems in the 

spacecraft. Escape systems in the Gemini were to be typical of 

current Gemini configurations. These decisions resulted in the 

need for additional technical studies to establish the basis for 

a selection of the Titan III configuration to be used for the MOL 

mission. 

I-1 
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1-2 

Accordingly, in June 1965, SSD authorized Martin to proceed 

with a 60-day technical study to further evaluate the use of larger 

SRMs on the Titan III (with and without transtage) and against the 

changed MOL mission requirements, The preliminary studies carried 

out in 1964 showed that incorporating larger SRMs in the Titan III 

system for the MOL mission was a growth option that substantially 

increased the payload capability into polar orbit and could be 

achieved with minimum change to the existing core structure for 

attaching the larger solids. The objective of minimum change to 

the core structure was obtained by selecting the thrust characteristics of the SRMs so that in-flight accelerations and airloads 

were no more severe than for the Titan IIIC with the present 5-

segment 120-in. diameter SRMs. Results obtained from the research 

and development programs for large SRMs at diameters of 120 and 

156 in. suggested the feasibility and practicability of procuring 

new, larger SRMs for the Titan III/MOL system with low technical 

risk. Payload performance capability was shown to be directly 

proportional to the total impulse derived from the SRM stage for 

motor sizes transportable by rail or road. Preliminary study 

results, however, did not fully explore the effects of the several 

varieties of larger SRMs that might be used for the MOL booster 

in terms of the specific modifications to the core structure for 

attachment, the flight control systems required for stability margins and load relief, the length of payload that could be carried, 

the technical risk of new SRM developments, SRM staging, the reliability of the system, the applicability of existing thrust vector 

control (TVC) systems, overall technical risk, development schedule 

requirements, etc. 

Therefore, the work reported in this document has as its primary purpose the technical assessment and comparative evaluation 

of using larger SRMs on the Titan III for the MOL mission. Three 

specific SRM options were studied. These were: 

1) 7 segment, 120-in, diameter (7 seg-120); 

2) 2 center segment, 156-in, diameter (2 CS-156); 

3) 3 center segment, 156-in, diameter (3 CS-156). 

Each of these options is evaluated for use with and without 

the transtage, making a total of six configurations studied. The 

7 seg-120 SRMs and the 3 CS-156 SRMs attach forward in Stage II 

creating local changes in the vibration in the forward compartments. The 2 CS-156 attaches at the present attachment location 

of the 5 seg-120 SRM at Vehicle Station 504 in Stage I. For the 

7 seg-120 configuration, a 15:1 expansion ratio is used for the 

LR-87 Stage I engines. 

6111111111rnia 
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1-3 

It is anticipated that the results of this 60-day study will 

contribute to selection of the configuration(s) that will be carried through the final definition phase for the Titan III/MOL 

booster. The effort covering detailed definition of the modifications required for crew safety and mission success is not included in this report. Additionally, the separate contractual 

effort to define the criteria for the interim launch capability 

(ILC) facility at WTR is not reported in this document, This 

study did not include consideration of larger diameter core stages 

for the Titan III or compare the relative cost/effectiveness of 

this system's growth option to the options using the larger SRMs 

for the MOL mission. 
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II, REFERENCE CONFIGURATIONS 

This chapter describes the six vehicle configurations studied 

during the period covered by this report, Presented are: (1) 

details of the core modifications to accept each performance improvement configuration; (2) a listing of crew safety/mission 

success mods assumed for the six configurations; (3) comparative 

data for the three SRMs under consideration; and (4) information 

on the payloads evaluated. 

The six basic configurations are listed in Table II-1. 

Table 1I-1 Reference Configurations 

Configuration 

No, SRM 

Stage I 

Engine 

Expansion 

Ratio Transtage 

Specified 

Payload 

Length 

(ft) 

1 7 seg-120 15:1 Yes 54.5 

2 7 seg-120 15:1 No 74.5 

3 2 CS-156 8:1 Yes 58.5 

4 2 CS-156 8:1 No 78.5 

5 3 CS-156 8:1 Yes 61.0 

6 3 CS-156 8:1 No 81,0 

Figure II-1 presents the general arrangements of the six configurations. 

A. CORE CONFIGURATIONS 

Each of the six performance improvement combinations required 

modification of the basic Titan IIIC core configuration, These 

modifications are described in the following subsections, A 

summary. uf the changes required on Titan IIIC core (Vehicle 11 is 

the reference) for each of the six configurations is listed in 

Tables 11-2 and 11-3. 
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11-2 

Certain other core modifications were required to add crew 

safety/mission success hardware. This class of modification was 

necessary so that the vehicles under study would more closely 

approximate the final MOL booster, which will have such additional 

equipment. 

1. Performance improvement Modifications, 7 seg-120 with Transtage 

a, Structures 

Stage III -- Equipment Module (Compartment 3A) (New) - The 

relocatien of antennas in this module is due to pattern 

interference by the longer SRMs. This performance change, 

along with the addition of redundant equipment on the 

trusses for crew safety, causes the module to grow 13.5 

in. in length, 

Stage III -- Propulsion Module (Compartment 3B) - Gages of 

several skin panels will be increased to meet higher panel 

flutter requirements, 

Stave II -- Compartment 2A New - The forward outriggers 

now attach to this skirt, necessitating the addition of 

new frames, heavier skins, and provisions for the SRM forward electrical interfaces. To increase flexibility, a 

bolted flange is added to the aft end of the skirt for 

attachment to the oxidizer tank, 

Stage II -- Oxidizer Tank - Since the oxidizer vent valve 

is moved from the skin to the forward dome (to avoid the 

high acoustical loading on the skin due to the long SRMs), 

brackets must be added to the dome to support the valve. 

A bolting flange is added to the forward end of the tank 

barrel, 

Stagfl::peat Shields - Due to the 15:1 expansion ratio 

engine, new, larger thrust chamber assembly covers (heat 

shields) are required, as are some modifications to the 

engine bell insulation panels, 

Other Modifications -- Core Insulation - Insulation must 

be modified in the vicinity of the SRM nose cone shock 

wave impingement on the core, and where the SRM staging 

rocket plumes impinge on the core, 

UNCLASSIFIED 
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11-5 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 

Other Modifications -- SRM Forward. Outriggers - The outriggers and fittings must be beefed up due to higher tension 1.oads created by thrust termination. The joint now 

existing on the core end of the outrigger will be redesigned to prevent imposition of bending loads on the 

staging stud, 

b, Ordnance 

No change. 

c. Propellant and Pressurization 

The Stage I pressurization levels might have to be changed 

to meet the NPSH requirements, as well as to meet any new 

specifications for the 15:1 engine. These changes might 

lead to higher tank top pressures. 

d, Flight Controls 

The flight control hardware will be modified to accommodate 

the longer burn time and environmental changes that result 

from the SRMs and the differences in load dynamics incurred 

with the heavier, longer payload capability. The modifications are: 

Provide an additional gain state and filter value for 

Stage 0 flight; 

Relocate the lateral acceleration sensing system (LASS) 

(load relief accelerometers) and Stage II rate gyros 

into the forward section of the transtage. 

Hydraulics 

A new or revised Stage I actuator (i.e., a new snubber) is 

required to accept the additional loads of the 15:1 engine, 

In addition, the redesigned Titan IIIB Stage II actuators 

must be used. The effect of the bending modes of the MOL/ 

Titan III booster, coupled with the natural frequency of 

the engine truss/actuator, could cause the actuator to go 

unstable. This new Titan IIIB unit is believed to be usable for the Titan III/MOL booster. 
f. Electrical 

Because of the increased length of the 7-segment SRM, the 

forward attach point and electrical interface connections 

and circuitry are relocated forward in Stage II, 

g, Malfunction Detection System (MDS) 

No change. 

h. Tracking and Flight Safety 

No change. 

i. Instrumentation 

No change. 

j. RF Systems 

The command control antennas in the transtage are moved 

to BL 0 to avoid the interference caused by the longer 

solids. 

The Stage II telemetry antenna system is deleted because 

of unacceptable vibration levels at its Titan III location. The Stage II telemetry transmitter uses the transtage antenna system. To accomplish this, a new fourinput multiplexer and a coaxial staging disconnect are 

required. 

Performance In rovement odifications 1 se -120 wi 

stage 

a. Structures 

Stage II -- Equipment Module (C artment 2A New Forward 

Skirt) - Relocation of antennas and umbilicals previously 

located in the transtage is required. Provisions must be 

added for SRM forward outrigger attachment (new frames) 

and for SRM forward electrical interfaces. New, more efficient equipment trusses will be used. This is possible 

since the transtage propellant tanks/engines do not intrude into this area. A bolting flange is added to the 

aft end of the structure. 

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11-8 

Tran 
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11-9 

Stage II -- Oxidizer Tank - Some as 7 seg-120 with transtage. 

Stage II -- Compartment 2B. - Add Glotrac antenna (skinmounted) at EL O. (This antenna is mounted on the attitude 

control motor fairing on the transtage on the Titan IIIC 

core.) 

Stage I -- Heat Shield - Some as 7 seg-120 with transtage. 

Other Modifications -- SRM Forward Outrigger- Same as 7 

seg-120 with transtage. 

b. Ordnance 

The payload staging hardware will require an increase in 

size due to higher tension loads without a transtage, 

This is a result of P increasing as a function of vehieq • 

cle station and the absence of axial load relief from the 

transtage. The staging nuts will be increased from 3/4 

to 1.0 in. diameter. The number of separation points 

will remain the same as will the interface dimensions. 

All Stage III items are deleted, 

c. Propellant and Pressurization 

Same as 7 seg-120 with transtage, but with all Stage III 

items deleted. 

d. Flight Controls 

Provide an additional gain state and filter value for 

Stage 0 flight (because of longer burn). Move all needed 

equipment into Stage II and delete Stage III equipment, 

e, Hydraulics 

Same as . 7 seg-120 with transtage, but delete Stage III 

hydraulics and actuators. 

f. Electrical 

Same as 7 seg-120 with transtage, but move needed 

equipment into Stage II and delete one IPS system 

because of transtage deletion. 
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II-10 

g. MDS 

No change, but move needed equipment into Stage 

h. Tracking and Flight Safety 

Delete the Stage II inadvertent separation destruct system 

(ISDS). Since Stage II is the last powered stage now, no 

ISDS is needed. In addition, all necessary tracking and 

flight safety equipment is moved into Stage 

i, Instrumentation 

Move all needed equipment into Stage II from the transtage. 

j. RF Systems 

Move all antennas to BL 0 in Stage II. Delete one telemetry antenna (reference Vehicle 11 has one antenna in the 

transtage and one in Stage II) and provide a new multiplexer so that all transmitters can share one antenna. 

Move the Glotrac antenna between Stage II tanks, since 

there is no further room for antennas' on BL 0 of the new 

Stage II equipment module. 

3 Performance Improvement Modifications, , 2 CS-156 with Transtage 

a. Structures 

Stage III -- Equipment Module (Compartment 3A) (New) - 

Same as 7 seg-120 with transtage. 

Stage III -- Propulsion Module (Compartment 3B) - Same as 

7 seg-120 with transtage, 

Stage II -- Compartment 2A - Skin gages are increased to 

meet panel flutter requirements and a bolting flange is 

added to the aft end of the skirt to provide flexibility. 

Stage II -- Oxidizer Tanks - A bolting flange is added to 

the forward end of the tank barrel, 

Stage II -- Compartment 2C - Provisions for the SRM electrical interfaces must be modified for more and/or larger 

connectors. The interface fairing must be modified accordingly. 
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Stage I -- Compartment lA - The frames in the areas of the 

SRM forward outrigger attachment must be beefed up to accept: higher loads. 

Stage I -- Fuel Tanks - Due to large differential tailoff 

thrusts on the 156-in. SRMs, the barrel, "K" frame, and 

longerons must be strengthened, 

Other Modifications -- Core Insulation - Same as 7 seg'-l20 

with transtage, 

Other Modifications SRM Forward Outriggers - Same as 

7 seg-120 with transtage, 

b. Ordnance 

No change. 

c. Propellant and Pressurization 

The Stage I pressurization levels might have to be increased to meet NPSH requirements. 

d. Flight Controls 

Same as 7 seg-120 with transtage. 

Hydraulics 

The same actuator redesign is required as was described 

for the 7 seg-120 with transtage. 

f. Electrical 

No change except for possible increase in the core/SRM 

electrical interface. 

g. MDS 

No change. 

h. Tracking and Flight Safety 

No change. 
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11-12 

Instrumentation 

No change. 

j. RF Systems 

Same as 7 seg-120 with transtage, except Stage II telemetry 

antennas are retained for this configuration only. 

4, Performance Improvement Modifications, 2 CS-156 without Transtage 

a. Structures 

Stage II -- Equipment Module (Compartment 2A) (New Forward 

Skirt) - Relocation of the antenna and umbilicals previously located in the transtage, is required. New, more 

efficient equipment trusses will be utilized. Size of 

the payload staging nut is increased from 3/4 to 1 in. 

diameter and a bolting flange is added to the aft end of 

the structure. 

Stage II -- Oxidizer Tank - Same as 2 CS-156 with transtage. 

Stage II -- Compartment 2B - Same as 7 seg-120 SRM without 

transtage, 

Stage II -- Compartment 2C - Same as 2 CS-156 with transtage. 

Stage I -- Compartment lA - Same as 2 CS-156 with transtage. 

Stage I -- Fuel Tank - Same as 2 CS-156 with transtage. 

Other Modifications --,Core Insulation - Same as 7 seg120 with transtage. 

Other Modifications -- SRM Forward Outrigger - Same as 

7 seg-120 with transtage, 

b. Ordnance 

Same as 7 seg-120 without transtage. 
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c. Propellant and Pressurization 

Same as 2 CS-156 with transtage, 

d. Flight Controls 

Same as 7 seg-120 without transtage. 

e. Hydraulics 

Same as 2 CS-156 with transtage plus deletion of Stage III 

hydraulics. 

f, Electrical 

Same as 2 CS-156 with transtage. 

g, MDS 

No change. 

h. Tracking and Flight Safety 

Same as 7 seg-120 without transtage, 

i. Instrumentation 

No change. 

. RF Systems 

Same as 7 seg-120 without transtage. 

Performance I rovement Modifications 

a. Structures 

CS-156 with Transta e 

Stage III -- Equipment Module (Compartment 3A) (New) -

Same as 7 seg-l20 with transtage. 

Stage III -- Propulsion Module (Compartment 3B) - Same as 

7 seg-120 with transtage. 

Stage II -- Compartment 2A (New) - Same as 7 seg-120 with 

transtage. 
UNCIfigmrn 

Stage II -- Oxidizer Tank - Same as.7 seg-120 with transtage. 

Stage I -- Fuel Tank - Same as 2 CS-156 with transtage. 

Other Modifications -- Core Insulation - Same as 7 seg120 with transtage. 

Other Modifications -- SRN Forward Outri ers - Same as 

7 seg-120 with transtage. 

b. Ordnance 

No change. 

c. Propellant and Pressurization 

Same as 7 seg-120 with transtage except that pressurization changes associated with the 15:1 engine are not required for this configuration. 

d. Flight Controls 

Same as 7 seg-l20 with transtage. 

e, Hydraulics 

Same as 2 CS-156 with transtage. 

f. Electrical 

Same as 7 seg-120 with transtage. 

g. MDS 

No change. 

' h. Tracking and Flight Safety 

No change. 

i, Instrumentation 

No change, 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

11-14 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

UNCIMTIED 

. RF Systems 

Same as 7 seg-120 with transtage. 

Performance Im rovement Modificatioue 3 CS 156 without Transtage 

a. Structures 

Stage II -- Equipment Module (Compartment 2A) (Ne For ard 

Skirt) - Same as 7 seg-120 without transtage. 

Stage II -- Oxidizer Tank - Same as 7 seg-120 with transtage. 

Stagp II -- ComRartment 2B - Same as 7 seg-120 without 

transtage. 

Stage I -- Fuel Tank - Same as 2 CS-156 SRM with transtage. 

Other Modifications -- Core Insulation - Same as 7 seg-120 

with transtage. 

Other Modifications -- SRM Forward Outriggers - Same as 

7 seg-120 with transtage. 

b. Ordnance 

Same as 7 seg-120 without transtage. 

c. Propellant and Pressurization 

Same as 3 CS-156 with transtage. 

d. Flight Controls 

Same as 7 seg-120 without transtage. 

e. Hydraulics 

Same as 2 CS-156 with transtage. 

f. Electrical 

Same as 7 seg-120 with transtage. 

MCI MIRED 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

g. MAS 

No change. 

h. Tracking and Flight Safety 

Same as 7 seg-120 without transtage. 

i. Instrumentation 

No change. 

j, RF Systems 

Same as 7 seg-120 without transtage. 

7. Assumed Changes for Crew Safety/Mission Success Improvement 

A summary of the crew safety/mission success changes to the 

Titan IIIC core (Vehicle 11) that were assumed for each of the 

six configurations are summarized in Tables 11-4 and 11-5, and 

are explained in more detail in the following paragraphs. 

All Configurations with Transtage 

1) Structures 

Stage III -- Equipment Module (Compartment 3A) (New) -

To accommodate the redundant equipment listed below, 

this module must be increased in length. A new equipment truss is needed, for mounting this equipment. 

Stage III -- Propulsion Module (Compartment 3B) - The 

engine truss is revised to accommodate the added 

hydraulic pump. 

Stage II -- Compartment 2B - The equipment truss in 

this compartment is changed because of the newly added 

equipment for crew safety/mission success. 

Stage I -- Oxidizer Tank - The aft dome of this tank 

is modified so that the redundant flight controls rate 

gyros may be mounted. 
111 :1 teJ t" 

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Sem as Configuration I. 

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Oxidizer Tank 

Sameas ConfigurationI. 

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1. Increase length cc loosea ll equip. 

meet. 

2, Equipmenttones changetoaccomo- dateall new hardware. 

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2 es. 2 2 ' .3 . 

3 

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. 

1 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

11-18 

2) Ordnance 

The Titan III/MOL interface staging ordnance is changed 

to provide redundant nuts in place of the present manifolded nut. The electrical interface circuitry must 

be revised to activate the ordnance on the payload 

side of the interface. 

3) Propellant and Pressurization 

No changes for crew safety/mission success for this 

configuration. 

4) Flight Controls 

The flight controls system has two parallel, redundant 

channels from sensors through the electronics for all 

phases of flight. Dual sensors provided are the Stage 

I and II rate gyros and a lateral acceleration sensing 

system (LASS). The autopilot used is an analog unit 

(similar to the present Titan III system) consisting 

of dual computer and adapter programers. Dual static 

inverters are used to power necessary portions of the 

flight controls system. 

For Stage 0 flight, an extra (reference) autopilot is 

added to the dual units along with additional Stage I. 

and Stage II rate gyros, LASS, and static inverters. 

5) Hydraulics 

Redundant hydraulics systems are provided in Stages 0, 

I, and III. Gemini launch vehicle-type redundant hydraulic actuators are used in Stage I and in Stage III 

(a single 'system remains in. use in Stage II), Separate power sources are used for each of the redundant 

systems; electrical sources are provided for Stage 0 

and III, and engine turbine power for Stage I. 

6) Electrical 

Redundant accessory power supplies (APS) and transient 

power supplies (TPS) are provided to be compatible with 

the redundant flight control systems. To provide maximum reliability through the critical Stage 0 portion 

of flight, a third source of discretes (i.e., a sequencer) is supplied to majority vote certain functions. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

11-19 

7) MDS 

The assumed MDS differs from that of a Titan III vehicle (Vehicle 11) by addition of a Stage 0 autopilot 

selector, a guidance and control selector (switches 

both guidance and flight controls), and a computer 

selector switch (to select which of the dual autopilots is controlling the vehicle), The MDS pressure switches of Titan III are changed to an analog 

type, usable for cockpit display. A delta pressure 

sensor is provided to monitor the difference in motor 

pressures between the two SRMs. 

8) Tracking and Flight Safety 

The ISDS enable/disable switches are made redundant. 

9) Instrumentation 

The number of measurements are increased to accommodate monitoring of the redundant equipment as well as 

to provide measurements for real-time ground displays 

for crew safety. 

10) RF Systems 

No changes for crew safety. 

11) Guidance 

The payload inertial guidance system (IGS) becomes a 

backup system for the booster IGS. 

b, All Configurations without Transtage 

1) Structures 

Stage II E•ui•ment Module Co artment 2A - The 

new Stage II forward skirt, required because of deletion of the transtage, is used to house the basic 

Titan III equipment; the increased equipment needed 

for crew safety/mission success affects the final 

length of this skirt. The new equipment truss is 

further changed because of the greater quantity of 

equipment. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

11-20 

Stage II -- Compartment 2B - Same as with-transtage 

configuration. 

Stage I -- Oxidizer Tank - Same as with-transtage 

configuration. 

2) Ordnance 

Same as with-transtage configuration. 

3) Propellant and Pressurization 

No changes for crew safety/mission success for this 

configuration. 

4) Flight Controls 

The flight controls system uses three parallel channels 

that are majority voted for the entire flight. The 

autopilots are the same as used in the with-transtage 

configuration with the transtage portions deleted. 

Three Stage I rate gyros, three Stage II rate gyros, 

three LASSs, and three static inverters are provided. 

5) Hydraulics 

Thg hydraulics system used in this configuration is 

the same as the with-transtage case, except that the 

Stage III system is not required. 

6) Electrical 

This system differs from that previously described 

(with transtage) by the addition of a third APS battery and the deletion of the second TPSpattery. 

7) MDS 

As the autopilots in this configuration are majority 

voted for all three stages, no malfunction sensing or 

switching is required in this area. This allows the 

deletion of the Stage 0 autopilot selector and the 

control portion of the guidance and control selector 

of the with-transtage case. The remainder of the system is as previously described. 

8) Tracking and Flight Safety 

Same as with-transtage configuration. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

11-21 SSD-CR-65-206 

9) Instrumentation 

Same as with-transtage configuration. 

10) RF Systems 

No changes for crew safety. 

11) Guidance 

Same as with-transtage configuration. 

Other Modifications 

Although not related to performance improvement or crew safety, 

other modifications are made to the core because the vehicle is 

flown from WTR. 

These include deletion of the command control receiver batteries and the Azusa rate beacon. The battery deletion is possible 

since the requirement for additional separate power sources for 

command receivers has been waived by WTR; thus, it is planned to 

supply these receivers from the APS and IPS buses for the MOL mission. 

Ground tracking stations for Azusa are located so that little 

data are received from the Titan III/MOL boost trajectory. Therefore, the Azusa can be deleted. 

Wei ht Summa 

The weight summary for the six configurations is given in 

Table 11-6. The vehicle liftoff weights are shown with the weights 

for the modifications previously described. 

Also shown are the delta increase in weight (above the minimum 

modification core) required to achieve a 99% launch probability 

(i.e., launch in full specification wind) with the specific payloads. 

A more detailed breakdown of the delta weights is presented 

in the following paragraphs for the minimum core modification numbers presented in Table 11-6, listing major items only. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 UNCLASSIFIED 

11-22 

Table 11-6 

,0 

Withou t Transtage 

(No. 6) 

0 

H ,o. 

o.,, 

0 

0 

m 

,, 

O 

riii 

0 

co 

...-■ = 

o 

0 

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rs) 

-4 

,-- 

el 

i 

-. 

, 

+ 

*This nweight inc ludes added equ ipment, w iring, etc, toperformt heMCI m issions. Structural modificat ions includedetc those minimum 

mod ificat ions not a ffected by changes inpay load configurat ion, 

tOnly structural beefup weights req u ired to achieve a997, launch probability w ith the specifiedpayloads are inc luded, 

■.t.N 

....,., ■ c.4 

,.... 

M 

a., 

04 

01 

4...1 

2,5370 100 2,911,952 

CO.....ft. 

0 V) 

11 

1-1 . 

0 

0 

H 0 0 c0 n 

Z .. 0 

.4 

.44 

M 

,C, 

•...t + 

7.1-' 

Without Transtage 

(No, 4) 

1,948,560 

2,296,337 

0 

un 

0 0 

0 

co r0 en 

0 

in fri I 

‘iin 

ir• H 

co 

With Transtage 

(No. 3) 

0 ‘,4 

1,948,560 

2,323,257 

0 

,7•■ 0 0 -I 

as 

.....i 

, + 

1) 

ao 

tn 4-. ,.o 

......, 

1,364,626 

1,713,263 

Il t`

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4..1 Er. 0 . ... .., 

4.1 

),1 c•-■ rir .0 iri r r 

.......1 H 0 

Booster Weight (lb) 

Core 

SRN 

Total Liftoff Weight 

Payload Specified 

Length(ft) 

Weight (lb) 

.e, 

[

771.47:17fo777.177:17— 

Beefup to Achieve 997, 

Launch Probability 

(1b)t 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 

11-23 

a. 7 seg-120 with Transtage 

Stage I -- Higher engine weight due to increased expansion ratio; larger 

thrust chamber cover plates; 

addition of redundant Stage 

actuator system 

Stage II -- Relocation of SRM forward outriggers; addition of Stage 0 

autopilot selector; addition of 

equipment 

+968 lb 

+374 lb 

Stage III -- Additional guidance equipment; 

addition of redundant Stage III 

actuator system; increased length 

of equipment module +325 lb 

+1667 lb 

b. 7 seg-120 without Transtage 

Stage I -- Same as 7 seg-120 with transtage +968 lb 

Stage II -- Replacement of oxidizer tank forward skirt with the Stage II 

equipment module; addition of 

equipment formerly located in 

transtage; relocation of SRM 

forward outriggers +1889 lb 

Stage III -- Removed -5647 lb 

-2790 lb 

c, 2 CS-I56 with Transtage 

Stage I -- Beefup of SRM forward outrigger 

attachments; addition of redundant Stage I actuator system; 

beefup of fuel tank due to high 

differential SRM tailoff loads 

Stage II -- Addition of Stage 0 autopilot 

selector; addition of environmental insulation; addition of 

equipment 

+510 lb 

+118 lb 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

11-24 

Stage III -- Additional guidance equipment; 

redundant Stage III actuator 

system; increased length of 

equipment module 

d. 2 CS-156 without Transtage 

+315 lb 

+943 lb 

Stage I -- Same as 2 CS-156 with Transtage +510 lb 

Stage II -- Replacement of oxidizer tank 

forward skirt with the Stage II 

equipment module; addition of 

equipment formerly located in 

transtage +1507 lb 

Stage III -- Removed -5647 lb 

-3630 lb 

3 CS-156 with Transtage 

Stage I -- Relocation of SRM forward outriggers; addition of Stage I redundant actuator system; beefup of fuel tank due to high 

differential SRM tailoff loads 

Stage II -- Same as 7 seg-120 with transtage 

Stage III -- Same as 7 seg-120 with transtage 

f. 3 CS-156 SRM without Transtage 

Stage I -- Same as 3 CS-156 with transtage 

Stage II -- Same as 7 seg-120 without transtage 

Stage III -- Removed 

+209 lb 

+374 lb 

+325 lb 

+908 lb 

+209 lb 

+1889 lb 

-5647 lb 

-3549 lb 

UNCLASSIFIED 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

11-25 

10. Environment 

A comparison of the vibration levels for each of the six configurations with those of Titan IIIC is shown in Table 11-7. An 

indication of the changes in frequency, as well as the predicted 

levels, are shown in the table. 

The number of components requiring requalification (DAT) because of the environmental changes are listed in the following 

tabulation. 

SRM Configuration With Transtage Without Transtage 

7 seg-120 48 49 

2 CS-156 24 24 

3 CS-156 50 52 
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RELEASE 1 JULY 2015 

11-26 

Table 11-7 Comparison Chart of Compartmental Vibration (grms) Levels 

Compartment Titan TIN 

7 seg-120 2 CS-156 3 CS-156 

With 

Transtage 

Without 

Transtage 

With 

Transtage 

Without 

Transtage 

With 

Transtage 

Without 

Transtage 

3A A/F 63 32.0 4 N/A 35.2 1 N/A 32.0 4 N/A 

3A Alt, Con. Sys None 28.0 N N/A 29,6 N N/A 28,0 N N/A 

3A Truss 16.7 10.8 4 N/A 9,1 1 N/A 10,8 4 N/A 

3B A/F 31.6 25.3 4 N/A 22,0 f N/A 25,3 f N/A 

33 Engine 26,9 26,9 4 N/A 26,9 f N/A 26.9 f N/A 

2A A/F (Tie 34,0 79,9 f 94,0 4 35,2 35.240-0- 79,9 4 94,0 4 

Point Fwd) 

2A A/F (Tie 34,0 87,0 4 87,04 35.2 -4-4- 35,2-4-0- 87,04 87,0 4 

Point Aft) 

2A Truss None N/A 17.7 N N/A 10.0 N N/A 17,7 N 

2A T/E None 56.4 N 56,4 N N/A N/A 56,4 N 56.4 N 

2B A/F 29.5 112.5 5 112.5 4 30.4-0-1, 112.5 + 30,4-4-4.- 112.5 4 

2B Truss 10,5 23.2 4 23.2 4 10.3-4-4- 10.3-4-4.- 23,2 f 23.2 4 

28 1/0 None 65.1 4 65,1 f N/A N/A 65.1 f 65.1 4 

2C'A/F 83,7 28.6 4 28.6 1 28.6 4 28.6 4 28.6 + 28.6 f 

2C Engine 68,0 68,0-4-4- 68.0-4-4- 68.0+4.- 68.0-4-0- 68.0-4-0- 68.0-4-0- 

IA A/F (at Tie 40,1 N/A N/A 34.2/ 34,2/ N/A N/A 

Point) 

1A A/F 188,0 33,0 4 33.0 4 None None 33,0 f 33,0 f 

1A T/D 59.4 34.0 4 34.0 4 67.8 4 67,8 4 34.0 + 34.0 4 

18 A/F 43.5 41.0-4--0- 41.0-4-0.- 58,1 4 58,1 4 58.1 f 58.1 4 

18 T/D 35.7 21,7 4 21.7 4 32.4 4 32.4 f 32.4 4 32,4 f 

IC A/F 45,0 38,3'1\ 38.3is, 40.O\ 40,0".. 40 , ON, 40 , 0 '‘'s, 

IC Engine 48.7 48.8-4-4.- 48 .8 -4-0- 48 .8 -4-0- 48 .8-4-0- 48.840-40- 48.8-'--4-' 

S1A, S2A A/F 52,4 52.4- - 52.4 -4-0- 52.4 -4-1-- 52.4-4-0- 52.4+0- 52.4-0-4- 

S lA , 82A Truss None 15.9 N 15.9 N 15,9 N 15,9 N 15.9 N 15,9 N 

SIB, 5213 A/F 46.2 46.2 46.24,1-0- 65.2 4 65.2 + 73,3 + 73,3 4 

TVC, Fwd 43,5 41,0-4-4- 41,0-4-0- 58.1 4 58.1 4 58.1 + 58.1 4 

TVC, Aft 45,0 38 , Plc, 38,3 40,0 40.0Nrs. 40,0N. 40,0*.. 

Above Titan IIIC Specification, Requalification Required. 

-4-0-Near Titan IIIC Specification, No Requalification Required. 

4, Below Titan IIIC Specification, No Requalification Required. 

"'Above and Right Frequency Shift, Requalification Required. 

'Above and Left Frequency Shift, Requalification Required. 

N New Specification, Qualification May be Required. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

11-27 

B. SRM CONFIGURATION 

Three SRM configurations were used as references for the six 

configurations. The 7 seg-120 is designed using the present 5 

seg-120 as the basis. The 2 CS-156, and 3 CS-156 motors represent 

new designs, but utilize the Titan IIIC-type thrust vector control (TVC) system. 

The three SRNs are described below. Table 11-8 lists the general characteristics of the various SRMs and Fig. 11-3 and 11-4 

present thrust and acceleration data, See Fig. 11-2 for typical 

configuration. 

1. Motors 

7 seg-120 - The 7 seg-120 configuration uses an eight-point 

star grain design in the forward and aft segments with a combination of conical or cylindrical shapes in the seven center segments. The forward end of each center segment is partially inhibited. The engine nozzle is conically shaped and canted 6 deg 

away from the core longitudinal axis, 

2 CS-156 and 3 CS-156 - The 2 and 3 CS-156 configurations use 

a 10-point star grain design in the forward segment with either 

a conical frustum or a combination of conical and cylindrical 

shapes in the center and/or aft segments. No inhibitor is used. 

The engine nozzle is partially submerged and is also canted 6 deg. 

2, T rust Vector Control (TVC) Systems 

TVC is obtained for the three configurations by pressurized 

N

2

0

4 

injection into the divergent portion of the engine nozzle. 

Each SRM configuration uses 24 injectant valves (six in each 

quadrant) canted 6 deg in the direction of the exhaust stream to 

produce a minimum thrust vector deflection of about 4 deg for 

any SRM thrust level condition. In addition, the TVC system can 

produce a minimum slew rate of the thrust vector equivalent to 

about 10 deg/sec. 

Hydraulic power supplied to operate the servo-controlled injectant valves is obtained from redundant electrically-driven 

pumps. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 Characteristics Table 11-8 

11-28 SSD-CR-65-206 

A in 

C..) 

in 

! Total ActionTime(sec) 124.3 138.0 155.0 

Throat Diameter (in.) 40.7 39.1 45.5 TVC Valves 24 24 24 Usable Inject Fluid (16) 6,248 9,988 1 1,050 

Length, Overall (in.) 1311.5 1146.5 1410.5 

Nozzle Type Conical Conical Conical 

(0% Submerged) (197. Submerged) (17% Submerged) ? 

Expansion Ratio 9.6 10.2 8.0 

Total Weight* (lb) 1,364,626 - 1,948,560 2,537,100 

Overall Mass Fraction 0.864 0.881 ' 0.888 

Total VacuumImpulse* (16-sec) 308 x 106 444 x 106 588 x 106 Maximum Sea Level Thrust* (lb) 2.97 x106 4.38 x 106 5.81 x1.116 

. , 

MEOPt (psi) 920 1 190 1190 

Liftoff Acceleration* (g) 1.74 1.84 , 1.99 

MaximumDynamic Pressure (16/£t2) 884 906 886 

Aeroheating Indicator (ft-lb/ft2) 93.0 x 106 97.5 x 106 86.6 x 106 

I Maximum Inflight Acceleration* (g) 3,08 2.96 3.00 

*Two SRMs. 

tMaximum expected operating pressure. 

(NI 

44:7 47 

02 

1,, Characteristic 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 

11-30 

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t. h tran

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 

0 

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0 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 

11-32 

The electronics to drive the TVC valves are located in Stage 0 

for the 7 seg-120 configuration. For both 156-in. designs, the 

TVG valves are driven directly from the core autopilot. 

3. Structure 

Each configuration contains a forward nose fairing and an aft 

skirt joined to the SRM segments (Fig. 11-2). The nose fairing 

encloses the head segment containing the ignitor and thrust termination ordnance, and contains the three electrical umbilicals, 

forward staging rockets (four for 7 seg-120, five for 2 and 3 CS156), two thrust termination ports, and the forward core vehicle 

attachment point. 

SRM segment cases are joined by a pin and clevis-type joint. 

The TVC system (7 seg,120 only), electrical and hydraulic power 

supplies, instrumentation packages, and the aft support truss are 

mounted on the SRM nozzle assembly inside the aft skirt. The aft 

skirt also contains the 4 aft separation rockets. 

The TVC pressurization and 14,04 tanks, the destruct raceway, 

and electrical raceway are attached external to the SRM segments. 

C. PAYLOAD CONFIGURATIONS 

The general configurations of the payloads used are shown in 

Fig. II-1, The Laboratory Module, of lengths shown, is 126 in, 

in diameter. The payloads used are topped by a Gemini B capsule. 

Lengths, weights, and centers of gravity (cg) are shown in 

Table 11-9. 
Gemini 

Capsule 

Overall 

Length 

Laboratory 

Module 

126-in, 

Dia —Titan III/MOL 

Interface 

cg 

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RELEASE 1 JULY 2015 

11-33 

Table II-9 Spacecraft Configuration 

Item 

7 eeg-120 2 CS-156 3 CS-120 

With 

Tranatage 

Without 

Transtage 

With 

Transtage 

Without 

Tranatage 

With 

Transtage 

Without 

Transtage 

Contract Payload Length (ft) 54.5 74.5 58.5 78.5 61,0 81.0 

Lengths Checked for Maximum 54.5 64,7 50,0 60,2 50.0 60,2 

Length Study (ft) 65.0 75.2 58.5 68,7 61,0 71.2 

75.0 85,2 65.0 75,2 72.0 82.2 

Payload cg each corresponds 29,6 27.1 27,1 

to maximum lengths above) (ft) 35.2 31.7 33.1 

40,6 35.1 39.2 

Payload Weight (lb) 28,000 28,000 33,000 33,000 42,000 42,000 

Spacecraft Geometry 
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III. PERFORMANCE AND TRAJECTORY CONSTRAINTS 

This chapter presents the payload capability, mission requirements, trajectory constraints, evaluation of the Gemini trajectory 

constraints, and trade studies that were performed. The payload 

capability of the basic Titan III core (with and without transtage) 

in conjunction with 7 segment 120-in. SRM (7 seg-120), 2 center 

segment 156-in. SRM (2 CS-156), and 3 center segment 156-in. SRM 

(3 CS-156) is presented for an 80-n-mi circular polar orbit 

[Western Test Range (WTR) launch] and for due east launches from 

the Eastern Test Range (ETR). The ETR performance capability is 

presented for a 100-n-mi injection using a park orbit ascent technique, or a direct injection ascent technique. High-altitude 

circular and elliptical orbit payload capability is presented for 

the ETR launches. 

A. PERFORMANCE TRADE STUDIES 

Design criteria were established to evaluate the 7 seg-120, 

2 CS-156, and 3 CS-156 SRMs. These criteria were as follows: 

1) Longitudinal load factor (T-A)/W within 5 sec after 

liftoff, larger than 1.6 g; 

2) Maximum (T-A)/W (Stage 0 phase), less than 3.2 g; 

3) Maximum dynamic pressure (qmax)) less than 900 Ibift

2

; 

4) Maximum differential thrust at Stage I engine ignition, 

less than 280,000 lb; 

5) Maximum aerodynamic heating indicator pV3 dt = 

95 x 106 ft-lb/ft2

, 

where: 

T = Vehicle centerline thrust (lb) 

A = Axial aerodynamic force (lb) 
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111-2 

= Vehicle weight (lb) 

/ 3) 

= Atmospheric density (slug/ ft / 

V = Relative velocity (fps) 

t = Burn time (sec). 

These criteria were established to assure satisfaction of vehicle and payload constraints, such as acceptable thrust-to-weight 

ratio at Liftoff to avoid excessive launch drift; specification 

allowable as a payload and vehicle constraint; maximum 

of qmax 

longitudinal load factor to satisfy core structural design criteria; 

and differential thrust at Stage I ignition to meet Stage I longeron 

design loads criteria. Thrust regressivity was not specified so 

the SRM manufacturer could exercise latitude in propellant grain 

design and internal motor ballistics. Core configuration data 

were given to each of the SRM manufacturers to allow them to tailor 

the SRM thrust, total impulse, mass fraction and mass decay rate 

to achieve optimum payload performance while meeting the constraint 

conditions. SRM manufacturers supplied data on the 7 seg-120, 

the 2 CS-156, and the 3 CS-l56 SRMs. Detailed SRM characteristics 

for the three configurations are presented in Table III-1. The 

thrust time histories are depicted in Fig. III-1. 

The 80-n-mi WTR nominal performance (with margin withheld) 

was obtained by using the SRM characteristics presented in Table 

III-1. This performance capability is shown on Fig. Note 

that payload weight is almost directly proportional to the SRM 

total impulse. 

Representative (T-A)/W (longitudinal load factor) time histories with tabulation of q (T-A)/W near liftoff, (T-A)/W (maximax, 

mum), and aerodynamic heating indicator are shown in Fig. 111-3. 

Several meetings were held between the SRM manufactures and the 

Martin Company to obtain the SRM characteristics (Table 

Fig, 111-1, and 111-2) that would satisfy the vehicle and trajectory constraints (Fig. 111-3) and yield the maximum payload 

capability. 
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111-4 

UNCLASSIFIED 

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Legend: 

2 CS-156 

-- 3 CS-156 

--- 7 seg-120 

Note: Curves for single S 

20 40 60 80 100 120 140 160 

Time (sec) 

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3 CS-156 886 86.6 X 106 3.00 

2 CS-156 906 97.5 X 106 2.96 

3 

Longitudinal Acceleration 

Limit During Stage 0 

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3 CS-156 

----2 CS-156 

1.„ 

Desired Liftoff 

Acceleration 

Note: 80-n mi WTR launch with optimum 

transtage propellant load. 

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111-6 

Time (sec) 

Fig. 111-3 Axial Acceleration Profile 
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B. PERFORMANCE CAPAB5L1TY 

The performance capability for the. selected MOL configuration 

vehicle was developed using the SRM configurations identified in 

the Performance Trade Studies (Section A of this chapter). WTR 

and ETR performance capability was determined for the 7 seg-120, 

2 CS-156 and 3 CS-156 SRM configurations with and without the 

transtage. The vehicle weights, SRM characteristics, liquid engine data, propellant loads, and SRM staging sequence are presented in detail in Section C of this chapter, Vehicle Characteristics, for the six configurations of interest. 

The pertinent ground rules that were used in the development 

of the reference trajectories are 

1) Initial azimuth from WTR, 182 deg; Iron ETR, 90 deg. 

2) A three-degree-of-freedom digital computer trajectory 

simulation program incorporating an oblate rotating 

earth model was used. 

3) The ascent trajectories were shaped by using an initial 

pitchover rate from 10 to 20 sec, a second pitch rate 

from 20 to 30 sec to drive the angle-of-attack to zero, 

a zero-lift pitch control through the remaining portion of Stage 0 operation, followed by a constant 

pitch rate from 150 sec to burnout of the final stage. 

4) The core engines were balanced at 60°F and 60°F propellants were used in all, liquid and solid stages. 

5) Stage I engine expansion ratio was modified from 8:1 

to 15:1 for the 7 seg-120 SRM configuration. 

6) TVC requirements (flow rates for control and dump 

schedule and resulting thrust augmentation were 

representative of those used on Titan IIIC. 

7) The Stage I engines were ignited during SRN tailoff 

when the axial acceleration decayed to 1.5 g. 

8) The SRMs were jettisoned 1 sec after burnout. 

9) The required velocity margin for minimum performance 

was estimated by root-sum-squared 2V4 of the ideal 

velocity of each stage. A burning time margin equivalent to this velocity was retained in the final stage. 

The nominal performance payload capability does not 

assume a propellant margin is retained in the final 

vehicle stage. 
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111-8 

1, Western Test Range 

Vehicle performance capability for launches from WTR are presented in Table 111-2. This table shows the performance for the 

various vehicles (both with and without transtage) in terms of 

weight of payload that can be injected into an 80-n-mi circular 

polar orbit. The transtage propellant load. was optimized to yield 

the maximum payload weight for each of the configurations using 

the transtage. In all cases, the required velocity margin (in 

terms of reserve propellant) remains in the final stage. 

Table 111-2 WTR Performance Summary, WTR Launch, 

Polar 80-n-mi Circular Orbit 

Pertinent trajectory parameter plots such as altitude-inertial 

velocity, altitude-time, and dynamic pressure-time are presented 

in Fig. 111-4, 111-5, and 111-6 respectively, for the vehicle configurations with transtage. Figures 111-7, 111-8, and 111-9 are 

for the configurations without transtage. 

The instantaneous impact points (IIP) are shown in Fig. III-10 

thru 111-15 for the six configurations. Step 1 and 2 impact points 

along with intermediate time points are presented on the figures. 
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111-14 

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SSD-CR-65-206 
5..(.7: Jettison SRMs (129 s 

\Impact (343 sec )A„)\ 

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impact (545 sec) 

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Fig. III-10 Instantaneous Impact Points, 7 seg-120 with Transtage 
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1 

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t (941.4 sec)'.". 

ISOCON40 

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"Impact (377.5 sec) 

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,Impact (579.1 sec) \ 

eye/ 

F 

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Fig. III-11 Instantaneous Impact Points, 2 CS-156 with Transtage 
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jettison SRMs (158.6 sec) \ 

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Fig. 111-12 Instantaneous Impact Points, 3 CS-156 with Transtage 
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111-18 

Fig. 111-13 Instantaneous Impact Points, 7 seg-120 without Transtage 
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Fig. 111-14 Instantaneous Impact Points, 2 CS-156 without Trans tags 
ettison SRMs (158. 

mpact (401.8 sec) 

/\ •ocoeeo is. 

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111-20 

Fig III-15 Instantaneous Impact Points, 3 CS-156 without Trans tags 
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The above data reflect the performance of the various vehicles 

without regard to any reentry abort constraints that may be imposed. However, Fig. 111-16 shows the reentry abort ceiling in 

coordinates of instantaneous apogee altitude and apogee velocity. 

The three trajectories shown are representative of the configurations (with transtage) considered. Except for a small region, 

they do not violate the limit. By trajectory shaping techniques, 

the flight path can be made to remain below the limit, and at the 

same time, better performance (payload in orbit) may be possible. 

The reentry abort ceiling, the temperature limit (Fig. 111-17), 

and the qu3 limit (Fig. 111-18) were taken from the Aerospace 

interoffice correspondence 65-2130-ELL-156, dated 1 June 1965, 

"Payload Constraints for Titan IIIC, 7-Segment SRM Studies." 

Figures 111-17 and 111-18 show the limits and the values obtained 

from the three trajectories. The temperature limit is only slightly 

violated while the airloads indicator qap contains a large margin. 
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111-22 

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111-25 

2, ETR Performance 

Circular and elliptical performance capability for due east 

launches was generated for the 7 seg-120, 2 CS-156, and 3 CS-156 

vehicles with transtage. An optimum transtage propellant burn 

was used in all cases. 

Two shaping techniques were used to develop the payload weight 

altitude charts: ascent using a park orbit, and ascent by direct 

injection into a transfer orbit. 

a. Ascent Using a Park Orbit 

For these cases, the vehicle waS direct injected into a 

100-n-mi park orbit. This was accomplished for low- and 

high-altitude payload capability. 

Low-Altitude Payload Capability - For the low-altitude 

missions the maximum payload weight is obtained by burning a portion of the transtage propellant to achieve the 

100-n-mi park orbit. The transtage is ignited a second 

time to achieve the transfer orbit perigee velocity. For 

the circular orbit capability, the transtage is reignited 

at apogee of the transfer ellipse to attain the required 

velocity for circularization. 

8i ^b-Alt ude Pa load Ca abilit - The high-altitude missions require the fully loaded transtage to be placed 

into the 100-n-mi park orbit if the maximum payload capability is to be obtained. This will require an early shutdown of Stage II, If Stage II was not shut down early a 

larger payload weight. could be placed into the 100-n-mi 

park orbit, but the transtage would then not have sufficient energy to deliver this payload to the required higher 

orbit, i,e., the vehicle is transtage-limited instead of 

booster-limited. 

b. Ascent by Direct Injection into Transfer Orbit 

This technique differs from the ascent by park orbit 

technique in that perigee will occur at the 100-n-mi altitude, i.e., the transfer orbit is obtained by burning 

out at an altitude of 100 n mi with a zero flight path 

angle and a sufficient overspeed velocity to obtain the 

required apogee altitudes. At apogee of the transfer 

ellipse (for the circular orbit cases) the transtage is 

again reignited to obtain the required circular velocity. 
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111-26 

If the mission does not require a park orbit this shaping 

technique should be used since the entire Stage II propellant load can be used to all, cases, resulting in a 

Larger payload capability for the high-altitude missions. 

A park orbit has the advantage of allowing the final inject point, or transfer maneuver, to be located at a particular location relative to the earth. 

For missions requiring the use of a parking orbit, the 

full capability of the lower stages (0, 1, II) can be used 

by use of higher altitude parking orbits. This will improve the payload capability above the altitude whore. 

Stage II is shutdown early. This payload gain, however, 

will not he competitive with that shown for the no-park 

orbit cases. 

The performance capability from ETR is presented in Fig. 

111-19 thru 111-39. 

Figures 111-19 thru 111-25, are for the 7 seg-120, Fig. 

111-26 thru 111-32 are for the 2 CS-156, and Fig. 111-33 

thru 111-39 are for the 3 CS-156 configurations. Each 

set of figures consists of the following: 

1) Payload weight to 100-n-mi park orbit as a function 

of the transtage propellant consumed. 

2) Payload weight-altitude chart for the park orbit cases. 

The payload weight-altitude charts show both the nominal 

performance and the minimum performance. Nominal performance assumes the vehicle performs precisely as predicted. 

3) Transtage propellant usage for the park orbit cases. 

These figures functionally show how the transtage propellant is used to attain first the park orbit, the 

transfer ellipse, and the final circular orbit. Of 

particular note is the vertical dashed line representing the point where the transtage is no longer used 

to attain the park orbit, and Stage II must then be 

shut down early for higher altitudes, 

4) Transtage propellant usage for elliptical final, orbits 

using a park orbit. Again the dashed line shows the 

attitude at which the transtage is no longer required 

for the park orbit and Stage II must be shut down 

early. 
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111-27 

5) Payload weight-altitude chart for the no-park orbit 

cases. 

6) Transtage propellant usage for the no-park orbit cases 

and final circular orbits. 

7) Transtage propellant usage for the no-park orbit cases 

and final elliptical orbits. 

3. Special Studies 

The previous ascent trajectories were shaped using an initial 

pitch rate from 20 to 30 sec to drive the angle of attack to zero, 

a zero-lift pitch control during Stage 0 operation, followed by 

a constant pitch rate from 150 sec to transtage burnout. 

By deviating from the zero-lift portion of flight (after 90 

sec of SRM operations) and using multiple upper stage pitch rates, 

a considerable payload increase can he obtained at the expense of 

other parameters. Figure 111-40 shows this payload increase for 

a matrix of two upper stage pitch rates 0, and 5c3}. This technique was not used for the MOL study becaue the aerodynamic heating indicator and maximum dynamic pressure constraint of 95 x 10

6 

ft-lb /ft

2 and 900 lby/ft

2 respectively would be violated. The 

aerodynamic heating values and maximum dynamic pressure values 

are presented in Fig. 111-41 and 111-42 for this shaping technique. 

This short study was intended as an example to show how the 

various trajectory parameters are varied by trajectory shaping 

Similarly, a more complete shaping study could be performed to 

reduce the rigid body loads. 
Transtage Prope llantConsumed during 

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Park Orbit PayloadCapability, 7 seg-120 

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111-28 

(qi) 711V-Fati peotAed. 
Altitude (nml) 

Note. 1. 100-n mi circular park orbi 

90-deg launch from ETR. 

0 5 10 15 20 25 30 35 

.Nominal Ierformance 

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111-29 

Payload Weight (103 lb) 

Fig. 111-20 Payload Altitude Profile, 7 seg-120, 100-n-mi Park Orbit 
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111-30 

Altitude of Circular Orbit (ft mi) 

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111-32 

0 5 10 15 20 30 35 

Payload Weight 10

3 

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Fig, 111-23 Payload Altitude Profile, 7 seg-120, 100-n-mi Perigee Overspeed 
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Altitude of Circular Orbit (nmi) 

Fig. 111-24 Transtage Propellant Usage, 7 seg-120, 100-n-mi Perigee toCircular Orbits 

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111-33 

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111-34 

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Apogee Altitude (nmi) 

Fig. 111-25 Transtage Propellant Usage, 7 seg-120, 100-n-mi Perigee toElliptical Orbits 

Jr 

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Fig. 111-26 Pa rk Orb it. Pay load Capability, 2 CS- 156 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-35 

SSD-CR-65-206 

en • 0 

(TO 30Tam poiXtqa 
Altitude (nmi) 

NRO APPROVED FOR 

RELEASE 1 JUL/111536 

Nominal Performance 

14-1 

Minimum Per 

0 10 20 30 40 

Payload Weight (103 lb) 

Fig. 111-27 Payload Altitude Profile, 2 CS-156, 100-n-mi Park Orbit 

50 
111-37 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

if} 0 H 

01 0) queitodoad a2u1suexa„ 

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nt Usage, 2 CS-156, 100-n-mi Park to Circular Orbits 
Propellant Usage, 2 CS-156, 100-n-mi Park to Elliptical Orbits 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

Apogee Altitude (nmi 

aueI Tadoad aSel sueaa, 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

10 

0 10 20 30 

Payload Weight (103 lb) 

Fig. 111-30 Payload Altitude Profile, 2 CS-156, 100-n-mi Perigee Overspeed 

A ltitude(nmi) 

104 

10 

40 50 

111-39 

.71 

Note: 1, No-park orbit. 

2. Burnout at 100 n mi and 'zero flight path 

angle. 

3. 90-deg launch from ETR. 

Nominal Performance 

Minimum Performance 

1 

'

iJ 
0 

U 

CC 

,41 CS-I56, l00-n-mi 

c1 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 UNCLASSIFIED 

111-40 

SSD-CR-65-206 

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to CircularOrbits 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-41 

(41 C

OI) luPiTadoid .a2cq simaj, 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 111-42 

0 

Transtage Propellant Consumed du

ring Boost 

Park Or

b it Pay lo

ad Capability, 

o 

CT 

(q1) 1ti2T at1 PuclAPcI 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

Note: 1. 100-n-mi p- :k 

2. 90-deg 17to TR. 

Nominal P2rfor . 

Minimum Parformaace 

10 

10 

i= 20 30 

( 0 

Payload Weight (10 10 3 

Fig. 111-34 Payload Altitude Profile, 3 CS-15f, 100-n-mi Park Ciric 

-Circular Or,its 
Transtage Propellant Usage, 3 CS-156, 100-n-mi Park to Circular Orbits 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-44 CR" ,206 

tr) 

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I 4.• 

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(9I 0 adO_Id aSE1S1-1141, 

Apogee Altitude (nm l) 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-45 F ig. 111-36 TranstagePrope llant Usage, 3 CS-156, 100-n-mi Park to Elliptical Orbits 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-46 

10 

Note: 1. No-park orbit. 

2. Burnout at 100 n mi and zero flight path angle. 

3. 90-deg launch from ETR. 

10 

10 

10 

0 10 20 30 40 50 60 

Payload Weight (10 lb) 

Fig. 111-37 Payload Altitude Profile, 3 CS-156, 100-n-mi 

Perigee Overspeed 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 , t 1 f. • ,.1 i! 7 1.

4: • :!..._1 71

; - r 1---1 1. 

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ab l:•::Prop

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ujidot 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-48 

(TT doad a`i3v1 euval 
0 C 0 

C 

‘.0 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

ITT -4 9 
UNCIASSIFB 

CR-65-206 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-50 

Fig. 111-

41 Trajectory Shaping, Aerody

namic He

ating, 

ken 

C 

c.4 

C 

c'.1 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 

-CR-65-206 

0 0 

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Cr) 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-52 

C. VEHICLE CHARACTERISTICS 

The vehicle characteristics outlined in this section were 

used in the performance and trajectory studies. 

1, SRM Staging Sequence 

The Step 0/1 staging sequence was initiated when the vehicle's 

longitudinal acceleration decreased to 1.5 g. At: this time Stage 

I engines were ignited. The SRMs were jettisoned. 1 sec after zero 

thrust in the thrust time table. 

Thrust-Time H. 

The thrust-time histories for the 7 see-120, 2 CS-I56, and 

3 CS-156 are shown in Fig. 111-43, 111-44, and 111-26, respectively. 

Time histories for Stage I w i th a i5:1 expansion ratio (used 

with. the 7 seg-(.20) and Stage I with an 8:1 expansion ratio (used 

with 2 C8-156 and 3 CS-I56) are given in Tables 111-3 and T11-4, 

respectively. The Stage II time histories (used in all configurations) are tabulated in Table 111-5. For all configurations 

16,000 lb of thrust and 305 sec specific impulse were used for 

Stage III, 

3. Liquid Propellant Inventory 

The liquid propellant inventory Is Labulated in Table 11.1-6, 

4. Thrust Vector Control (TVC) Fiul 1 Inventory 

The following TVC fluid loads were used for performaned calculations; 

7 seg-120 Loaded, 29,500 lb, 

Expended, 21,000 lb; 

2 CS-156 Loaded, 33,000 lb, 

Expended, 21,000 lb; 

3 CS-158 Loaded, 38,410 lb, 

Expended, 30,090 lb. 

These loads were not used for TVC studies. 
r 

Note: Curve for single SRM. 

20 40 60 80 100 120 140 

Time (see) 

o 

(i) 

$4 

SSD-CR-65-206

H 

0

0 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-53 

Fig. 111-43 Thrust-Time History, 7 seg-120 
100 120 2O 40 60 80 

Time (sec) 

Li 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-54 CR 63-206 

Fig, 111-44 Thrust-Time History, 2 CS-156 

Thrust (1

06 lb) 

140 160 

Note: Curve for single SRM. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

S8D-CR-65-206 111-55 

Note: Curve for single SRM, 

20 40 60 80 100 120 140 160 

Time (sec) 

Fig. 111-45 Thrust-Time History, 3 CS-156 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-56 

Table 111-3 Stage I Class Nominal, Expansion Ratio, 

15:1; Used with 7 seg-120 

Acceptance Test and Launch at 60°F 

Time from Total Total Flow 

87FS1 (sec) Thrust (lb)* Rate (1b/sec 

526,730 

527,000 

527,400 

527,660 

529,100 

531,000 

532,060 

532,060 

532,060 

532,060 

532,060 

532,060 

1761.6 

1762.5 

1763.9 

1764.7 

1769.6 

1775.9 

1779.5 

1779.5 

1779.5 

1779.5 

1779.5 

1779.5 

6 

10 

16 

20 

35 

55 

75 

90 

110 

125 

140 

150 

*Chamber centerline including gas generator. 

tAutogenous flow excluded. 

Cant angle: 2 deg. Exit area: 5628 in. 
Acceptance Test and Launch at 60°F 

Total 

Thrust (lb)* 

Total Flow 

Rate (16/sec)t 

479,417 1668.2 

476,713 1658.6 

477,006 1659.0 

477,348 1660.2 

478,442 1664.0 

480,760 1672,2 

481,704 1675.4 

482,088 1677.0 

482,190 1677.8 

482,190 1677.8 

482,190 1677.8 

482,190 1677.8 

Time from 

87FS

I 

(sec) 

6 

10 

16 

20 

35 

55 

75 

90 

110 

125 

140 

150 

*Chamber centerline including gas generator. 

tAutogenous flow excluded. 

Cant angle: 2 deg. Exit area: 3023.6 in. 

5-206 111-57 

Table 111-4 Stage I Class Nominal, Expansion Ratio, 

8:1; Used with 2 CS-I56 and 3 CS-156 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-58 

Table III-5 Stage II Class Nominal, Expansion Ratio, 

49.2:1; Used with All Configurations 

Acceptance Test and Launch at 60°F 

Time from 

91F8 (sec) 

1 

Total 

Thrust (lb)* 

Total Flow 

Rate (1b/sec)t 

6 95,971 306.1 

10 97,038 309.6 

16 98,724 315.0 

20 99,338 317.0 

35 100,643 321.1 

55 101,452 323.8 

70 101,759 324.8 

100 102,154 326.1 

120 102,157 326.2 

150 102,199 326.4 

180 102,201 326.5 

205 101,597 324.4 

*Chamber centerline including roll nozzle. 

tAutogenous flow excluded. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-59 

SSD-CR-65-206 

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01 

r-1 

rA+ 

0) 

.± 

it) 

0 CO 3-40 ,0 , t X 01 

,C1 3-40 01 

. 

.....1. ,e) 

...,1' 3/1 

N N 

A, 

N '0 

,-+ , 

04 0) ■1 • 

0 0 

CO 0 

...* 0 

111 0-1 

w 

04 

Nomina l Va lues at 60*F 

Loaded Propellant Expended be fore Stage 0 Liftoff 

Propellant Aboardat Liftoff Propellant Expended as Bleeds or Leakage 

During Stage 0 During Stage 

During Stage II 

Total Nonusable Propellant (Trapped, Vapor Retained) 

Mean Outage 

Propellant Consumed during Stage Operat ion 

Start Steady State ShutdownandTa iloff 

Pressuriza tion System Inert Gas Mean Outage Percentage (Based on Loaded Propellant ) Max imum Outage Percentage (Based on Loaded Propellant) 

...--, 

0 

r-I 

30 

0.1 u to 

a 

co 4., 

r%'5(3 

1-+ 

8 5 R 

0 X 1). 2 

Tab le 111-6 Liquid Propellant Inventory 
NRO APPROED FOR 

RELEASE 1 JULY 2015 

111-60 

5. Step Performance Weight Summaries 

The performance weight summaries of the three configurations, 

with and without transtage are given by step in Tables 111-7 thru 

111-12. 

6. Axial Force Coefficient 

The Stage 0 axial force coefficients for the 7 seg-120, 2 

CS-156, and 3 CS-156 (with and without transtage) are given in 

Fig. 111-46, 111-47, and 111-48, respectively. Stage I axial 

force coefficients are given for all configurations in Fig. 111-49. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

III-61 

SRMConfiguration 

3 CS-156 7 seg - 120 

287,790 180,758 

8,320 8,500 

296,190 189,258 

(2,240,990) (1,175,368) 

2,208,900 1,142,648 

30,090 21,000 

2,000 11,720 

2,537,100 1,364,626 

*Step 0 inc lude s two SRMs. 

,.0 

Lr) 

,-.4 

I 

C...) 

N 

.,.., 0 0 0 0 0 0 0 0 

'.0 0 s.0 0 0 0 0 

,..0 

el 0 (r) 0 C> 0 0 u'1 

.., ,s .. " " 4, . . 

01 N in ref 0 ri CN1 00 

(4-■ r.-1 .4 0 00 N -...1" 

N N 1--, 

,,C) (7, 

r-4 r-4 e-3 

,....^ 

Item 

Weight Empty 

Dry Weight 

Residuals 

TVC fluid trapped TVC fluid reverse 

Burnout We ight 

Flight Expendables 

Propellant 

TVC Fluid for Control and Dump Schedule 

Inert Material 

Step Weight at Vehicle Liftoff 

Table 111-7 Step 0* Performance Weight Summarywith andwithout Transtage 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-62 

al) 

14) 

0 

stt $.4 

E-1 

4.1 

0 

0 

0)3 

0 

Ca. Taole 111-8 

c.": 

r-1 

I 

a? 

Soa 

LI-, In 

L.0 ln 

•••:f tr) 

-4' 

p... 

CO in Co 

r. (r) --,t 

C r.4 if) 

In ,--i 

r4 

N- 

.--. 

-,± 

(N 

--.4 

N. 

‘-`,1 

...... 

- -+ ,---. (N Cl) 0 0-,0' c.4 -4 

1'. CO CO ...t In IN ,..... c"-.. 

.„-± (3, 0 C"4 C'l IN 

•,,O In Ln 

....4 In In 

N CV 

IN ..0 

In 

"..# 

IN 

N. 

C...1 

00 

In 

IN 

N. 

IN 

'---. 

CX 

0 

.e- 

..) 

0 ...-■ 

/.... 'D N M In In Co r..... ...„-j- *--i co r-4 IN CO 0 <7, a+ N ,...-4 0' `,C In 

a In ul 0 cn -,1- ---.' IN r---. In CO co -d In c•4 r-1 N. In •-...t 

CC r..-i '''.... In In ,...4 In CV r", 0, 0 NI 111 N N. " 

C6.4 Lo C.1 ,...# 4-4 

I 

In In In -. - 

a • r-i ,--i ,.....■ it) V". r--- NC 

‘.../ In 

N N 

`.-.' 

IN IN 

Z 

C:4 Cl) 

N re) 0 In 

0 In co 

..0 In 4 

N 

,--4 

..d. 

N 

r-4 

r•-• 

In •-■ N CO 0 0, 0' IN -4 

,--t co co -,1" cn (NI ,--, N• 

-....t ,.0 

a., 

0 0 

.- C) In Lel r....1 In Cs1 0 cn 0 N t''', N C7, 0 

....i. ■■ . r-1 ,..0 In In -a N 

0 r-t .-1 t-4 in in 

N Csa 

r•-•• 

Ca 

N, 

hi 

N

"......" 

' 

...**.., 

■---1 

..-I 

17.1 0 0 

C.) 0 -1-} 

•e4 • -1 4..] 

4.../ 4-1 al 

10 0:1 (14 0 ,1.4 

Cl )-4 :-.1 al 444&4-4 

G <1.% 0.) -•-.." 0 4-4 

•r•-■ 4-) 0. 0., 4.4 0 

(ti G 0 0 a) 4.4 4.4 

"J <14 S... -_-44-4 

Cl "-.1 0 1-1 0 

1.4 .1.1 CA r-1 

a/ CX 0 0 0 al 

E 

(1) 

.4.4 LA 

00 

O. a) 00 00 e.. 

CA 0 4.3 )--4 al al {-.) 

,.. al 

(..) -.1 

(CI 0. a) ). 0 .-I 4.444) ,-.1 0 4-3 

,--4 0 ,--d 04- M (1) at L"1) Cl) L. -0 4-1 ' 

r.4 > -0 r-4 04 on a) 1.1 0 CO 

CL) 

0 0 tz u... 

0 Cl *-1 0 }... n) 'Cl .0 

.."0 .1-■ a) )4 CI (1) G.) r- 

...i cl c‘z,...0 (1, cu ,.4 

> .0 

...J -0 

---I 

Cl 

.., ,v, 0 (31 ...0 0) .1..) 0 C.) 1--1 r-1 0 It (1.3 

7-, Cr: a.a 00 00 01-. Cl) 1. 0 CO CO C..) 3J ca. 

I., .>, • -0 p _u •,-4 x 1 CL. 4-4 ...., ,-.1 Cl) 

0 a 0 ,-4 a) r./..." ›, 0 a, a) a) 4-1 £ a:l ...) 

ri 4.-) 11 t'l CL Ca 1..1 ,4 ,.... 'a a-) '0 4-1 0 CX Cl) CO CI) 

(.1 0 CO 0 0. G 0 0 CO 4.) 0 ...1 -ci -,-4 .r.i .'- 0 .r.4 a) 

E 

4 ..) CO 

-0 t--1 

--00 0 

0 •-.4 )4 a/ 

til H 

,..4 

1. 

0 

VI 

Cll 

al 

4-1 

4-1 

tt 

43 ,0 al tt ;.1 00 00 ). 

0 00 l..1 4-1 ,.. 00 0 Z ,-0 

C) '14 CO GO to 1-4 [,x) CO H 

0 

--4" 

CO 

■-.4 

al 

• e-1 

0003 

- 4 0-i 

a) >,, c4 a ■--1 

P 1-1, 

0 CO 

r• ,---I 

s--■ L...

0-0 

<11 1.r2: 

...) 

-,-i 

0 4-4 4-) G 

::.., 

G4 CO CO t--1 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-63 

Table 111-9 Step 1 Performance Weight Summary without Transtage 

SRM Configuration 

C.) 

CNI 

,-t 

i 

tt 

4) 

0 N.. 

0 01 

...1- if) 

in in 

. 

4..t 

M VI 00 

cr, in „it 

0 ■-•4 trl 

. . 

AI) r-d 

,-,4 

r-- 

h.-4 

-...t 

CY 

,--.4 

r--. 

c.1 

i 

.-, 

,.0 4-1 N 00 0 

0, 0, (NI ,...I 

,1- 00 CO-.....t cn 8,4 *--i r--- 

in LT 0 c--4 on rg 

.. . . 

4) fl tcs, 

4.-4 v'') ,r) 8,1 8,1 

,......., 

N. \c, 

c‘l In 

. 

CA 1,- 8,4 

re) 

rn 

to 

. 

CV 4,-, C,J 

V, 

Ln 

r-d 

1 

CO 

0 

CC) 

CV rr'd 8,4 in 

CO di'd 

n 

O") 

,-.1 VI 00 

N. m ...1. 

Ccl rd-1 in 

n n 

....1. rd-1 

,....1 

r'... 

,..-1 

.4 

c',4 

4-4 

N. r■i 

...--.. 

00 ).-1 NI CO 0 C' 0`, N4 ....I 8,4 CO 00......1- C CN1 rd-1 r'dd. 

00 0, 0 8 V cn CV 

ds rd n 

u") I.C1 in 

,--1 in in 

CV CV 

‘...." 

Cr, V0 

0 

00 

n 

,-.I 

N 

CV 

In 

r-d 

CO 

n 

,...-i 

r-. 

CV 

\0 

in 

4-1 

I 

Li) 

U 

N 

rd, m 

N. in 

0 trl 

. 

...1 

0 in 00 

Cc) VI .1' ,CY 4-, )r) 

. 

...1. ,--4 

,-..1 

r.... 

dd-d 

.../d 

Cd 

1-4 

N. 

CA 

rd-d, 

Cc) r-dd CV CO 0 Ors 0`■ CV r-dd 

CO CO 00 .4' 01 

C,1 .4 N. 

CY ON 0 CA 

c,n 8.1 

. . . 

Y.0 If1 in 

..-4 U trt 

CV CV 

'v." 

...../. ‘.0 ,,J0 

0 

. 

c,4 

N. 8,1 

0 

4"-- 

0 

. 

C8 

N. 

CV 

4) 

4.4 

4-1 

›, 

.1.nd 

04 dnd 

0 4.4 

44 

4) 

dl) 0.) 

Xrd4 

CIO 4) 

4.4 E-4 

-0 

4) 

C 

.44 

(ti 

4-, 

0 

P6 

Z 0 

84 0. 

.4 

(4 

r-di 

4) 

O. 'V el) 

0 Z 00 

0 0 

a, 4-+ 

"0 C 

4.1 rd.di Qj 0 

,40 04 CI 

8.0 0 C:), C 

'd^d "0 0 to 

dld dr4 Sr 1.) 

0E4 Z 

a) 

›,t r4 

4.1 

Z 

RI 

C.J 

dd-I 

P 

XI 

hi 

4) 

0 

413 

a) 

0 

el3 

0 

4) 

P4 

4-4 

,-.1 

ca 

4-1 

4...3 

ra 

0 

CO 

..-■ 

.-4 •,-I 

0 is )4) 

0 0 1.4 

._4 4,1 4-1 

4-, 4.4 44 

44 44 0 

44 14 Al 

4) a`-' 

4-, 04 Cl, 

C 0 0 41) 

a) )4 

r-di 0 dd.-4 C 

r-i 4.-1 ti) 

W (1) <V 0 

04 44 CO b.0 .-1 

to 0 44, h.-4 40 44 0 

a) 

,-4 ar,rocv(vwc.,o;-, 

.M 1--I a, 00 al 

al 8J r.-1 0 ).. "0 

'0 4.4 8J S-.. RI 8.) tea) 8) E 

didddd C ctf 

0, 0,4 .0 a) al 0 

X 41 a.) 0 C.) 4-4 4-4 .0 

an at (/) $.4 Z CO tdO (...) 

.,-4 X I C.14 3 s..4 

0) 44 ›-, 0 4) tl, a) 4.4 -0 4.J -0 ..W Z 0 C8 

4.1 0 3.1 .d..d • el • •-4 drdi 0 

AJ X 4/ 0 0 C to 410 1.4 

0 013 diddl dd-d X 00 C C ,.0 

0 

,4 01 u0 to t--i 4.1 48 E-1 

Z ,--1 

)4 1.4.J 

444 

)4

,4 44 

04-, 

4-, 0 

LH 4-, 

.44 1.1-4 

1-di dr-i 

I-I 

4) 

.-t a) 

U $-.4 

..a ti-, 

a) at 

> .CI •LI I/ 

88 a) 

4) 

4.1 ed-d 

.0 CO 

8113 

ddddi 4,) 

4) C 

3 • d-4 

00 

0. 0 

a) f...1 

4.4 

X 

08 

-,-1 

ai 

00 

CO 

4-4 

C)) 

4-4 

tiS 

-,-1 

4.) ',-1 

a 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-64 Table III-10 Step 2 Performance We ight Summary with Transtage 

SRM Configuration 

0 

N 

in4 

by 

N U) 

N 

N ‘-r:, 

VI CT 0.1 ,JD 

Ul 

re) i 411 

01 \0 ,I• N 

‘,0 

r-4 

r.4 

4r1 a) 

/.01 

1 

0 0, u• r r, 00 (' ‘0 a, CO 

c1 ..,1" N a' ON Lri 

N 01 CO r-i ,-4 

N. N-• .0 

s.o .0 

...., 

ON N r, 

in 

...t 

r.-• 

H a) 

in 

--,t N., 

,0 

Ln 

v-41 

C/7 

C.) 

rn 

N. 421 oa a' 

ON VC, 

4r1 

Cfl ,t on 

co o ,-4 ,0 sst N 

l0 

1 

,--4 

,--f 

1.4-1 00 

irt 

1 

...-, 

0 ON tr1 in N CO 01 10 01 Cop 

LO 4 N a' cr, un 

N VI CO rni H 

N.. N.. %.0 

..0 s.0 

,..., 

os N 

0 

...0 

...1- r, 

r44 

44-1 

s.0 

n 

--1- N. 

,..0 

111 

r.1 

1 

Cr) 

N 

N 

,..0 

N ON 

LO LO 

n 

to 

c0 may in 

1-4 C) r.'1 4,-) -1r N 

n 

LO 

i-4 

r-4 

101 

0 

00 

1isi 

I 

..--. 

tf) ON In 4-e) N CO 4

,1 \D 0', CO 

CT ....t N Cr, ON to 

00 4,-) 00 1

-4 r

-1 

n n n 

LO N LO 

"...." 

...,e N 

..t 

N 

...1 

'0 

...t 

04 

,../' 

r-, 

41.) 

4.3 I-. 

› 

4.4 ›, Cl. )... 

5 4,.., 

rzl cl) 

5 

siss/ C.) 

X ,--1 

60 a) •,-4 E-i 

_co 

.0 

U 

(1) 

4.1 

4/ 

CO 

CO Q. 

ani Co ri > 

OS 

fa, IV al 

0 G 00 

P 03 

c1:1 

P4 4.1 

 G 

U s-4 (U 0 

4 to ia. 

b0 G Cl. G 

.,-; -a to fri 

0.) •e-1 0 

44) 

E-4 Z 

a.) 

›, CC s4 

0 

0 

al 

U 

-,-4 

F.4 

,.0 

..3 

(0 

tef 0 

Ci) 

i4 

0 

to 

0) 

34 

p., 

0.1 

al .'.1 

;-1 

U 

4-1 

(7) 

> 

ar'l 4.) 

 CO 

r-1 

4 <4 

04 

0 

0 0 0 

0 0 .r. 

..-1 .r-1 4-1 

44 4.1 Ct 

03 0 

-0 i.. CD 

01 0.1 

.1.1 CL Rn. 0 

0: 0 

100 

01 H 

rs4 0 t-t H 

0- 0 r.4 bl ) 0 0 tuti 

CO 0 141 s..41 CO CO SC03 

01 14 0 H •,-4 1.1 4.1 

4-i 

rni 04 03 a) 44 a) a) Cl) Cl) 

XI rn.I at 0) 00 

CV 0) r...1 0 U 4. .0 '0 'U 

'a 4-/ 0) 3-I (V 03 CO (1) CU 

4-1 0 0 0. O. Z ,..0 <0 0,) 0 

4 (1) .t..i 0 u H H 1..4 

OD C4 cr) s.4 p a) P A PA 

.1-1 X 1 

a■ > 1... 

11) 1.r.1 ›, C • 01 I:0 C) 0 U 

3 "1:1 4.1 "C1 1.1 4.) Z Z 0 

4.1 (0 4,1 Ct -I-1 •ri .r-1 ,--1 

4..) ..:::: cu CC 0 

■•••4 CC ba tZ tO 

0 tk) 4-) 4-) .0 ,o o0 0 0 0 

0 . el Cf) CO tr2 <4 1--1 44 4.1 fal 

Z r--; 

, 44 

03 

4.4 14-1 

4-4 

4.1 0 

44 .1.J 

...I li... 

i.-4 ..-4 

s444 CU 

U 14 .,..{ 0 

,C 4-4 

0) 40 >. .0 

4-) '0 

at 03 

a) 

4-1 r•-■ -0 00 

00 ,s-i co 

<11) ,...■ 

'3 se.' 

CO 

CL, C 

co tx.1 

4..J 

cr:, 

4-1 .,C 

00 ,-; 

(0 

o, 

a) 

4-J 

V) r-s 

cti 

.,-4 

-0 .,I 

CC 1-4 
a30 r, 

‘.0 8 2 C•4 

N. 00 

■D 

CO Cr, h 

r-1 

N. 00 

00 c) r■ C 

Ln N. CY, CR V) 

1 CO 

00 N.'0 

'0 '0 

N 

013 a) 

U) 

- 111 N 

'0 

r-- 

0 

CO 

4-1 

0 

0 

Li 

'0 

tr) 

U 

r--4 tr.; OD c•-1 Q\ LfC cn N. CO C) '.0 C N (".1 

Cr, in ..„1. N. Cs 01 If) 

r- 0.1 co 

CO N. ‘.0 

‘.0 N. 

4) 0\ It1 N- CO 01 ,0 0"■ 

cf, to 

c0 r-o 

Lrl N 

r-- 

0 

N. 

00 Lf1 

‘.0 N. N. 

0 0 00 

-,-4 -,-1 

4.1 4,-) 

all 0 4-4 

44 1.4 '1-19.4 

W0.1 04.4 

.-■ 0.0. 4., 0 

G 00 4-4 4.1 

CO •".. 4-1 

1.4 0-4 ,..-4 •,-.1 

,---4 ,.., •-J 

0) G 0)0) 

ca. CO,-4 0000 0.7 

0 C 4-.1 r--t co CO Ca Q1, 

W I, 

r--. Rd 

0 ,--■ 

cl CU 

-,-4 

I, 

4 1 :-/ 

W G1 0/1 LI-, 

0 .0 

.rs .-( C a . . a) tO W 00 

CO a) r--4 04.1 i.4 '-o ^a 4 .-.4 

CU 

4.) G c0 ca. 43. Z 4 0) W Ca0 cj ,--1 r-I W 

a) Q. 

.,i ya i A., > "... 4.1 "-I a) 

0.) 41 >, 0 ...-I 0) 0) cll 4 01 4.) 

.300 ti) tn 

4.) CO $-■ 4.) CS: ..-4 •,-4 -,-i 4-4 CV 

W G ,--4 

0 00 4.J 4., 4 4 00 0 0 -,-1 co 

00 -,-4 0.0 4.1 

S.4 C.,... 0)z0 . el 

44 0 

pea cn 4-1 

'0 

UC 

r 

crt ■0 a, try 00 

00 a r-4 LC' 

\O r 

Retained 

▪ 4.4 

O 0 MS sa. 

1-1 CO 

a) 

Ck• 

▪ 

a) 

O 0 00 

#••■ CO 

fa. ,1-1 

›, 4.1 13 

g 4 ca 

t.T.1 000 a. a 

'V 01 c0 

•,-+ 

X U) E-' ti) a) 

• E-1 

;-■ 

Lubricant 

Pressure Gas 

Ablative Material 

Ta

b le III-11 

a) 

60 

4-1 

0 

CA 

00 

•■-1 

;3 

CU 

0 

4-1 

N 

Ca. 

U) 

4.) 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-65 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-66 

SRM Configuration 

0 N 

r-i 

CG 

a; 

0--) 

,...4 a 

.0 in 

-1- 

4 

r-i .../* N -.I-r- in 0 

0 r-4 S. ,.0 m 

N I 

trl 

en m N Li') 00 cn Ln 0 

tn ,0 CO (--) N m 

el 4-4Cr ,--4 

tr) m N 

N N 

.10 CT, 

,./' 

CO 

N 

4.0 

In 

;-4 

in 

0 

0 (-1 r-i a, 

4 in 

....t 

i 

I."( ...,t N .....t r- tr 0 

0 ,--I 1-... '.0 el 

N 1 

tf) 

M el N u CO el In 0 

el '.C) 00 (-4 N Cr) 

m r(4 Cr, r-4 

VS t",1 N 

CV N 

•,...4 

q:, ON 

`NI. 

CO 

N 

)40 

Ln 

,--1 

4 

UT 

U 

N 

in n-I 

..0 a■ 

in Lel 

at 

,..0 ..„1- N ist r- in 0 

WI ,-1 l'••■ ..0 m 

o.4 I 

') 

Lel 

..4"). 

00 M N in 00 el in 0 

N N el 

N r-i 0N r--4 

or .4 e, 

In rn N 

N N ....4 

n-4 

in 

....t 

CC 

N 

E (1) 

4..) 

1--1 

>, 

4-) ›N cl. $4 4.S 

41 44 

E 

4-) 4.4 

4 ,1 

0/ 01 

(r4 E-I 

A). 

4-s 

.4-4 

40 

r-i 

,-4 

(1) ci. 

0 

••■ 0 

"0 0 

cu t-I 

(a 4.4 M 

4-1 a 

a 0 a 

R' a.) G 

CO 

-U 4 4.1 r.-4 r-4 

CO '.0 0 a:1 

Co a, C 0 ',"( 

(-4 0 4,4 C„„) 4-1 

r-I > ),(4 0 

(1) th 4J CO 4-) 

43. "c) (1) to 1; (.7 ra 

0 Z at) o <a t--4 z 

14 Ca tO 

324 4.4 (1) 4-4 4-I (14 

-1-) T-1 0 8 ,.., 

4 0 04 01 (0 "0 4..) 

(30 0 04 0 (A (r-I 41-1 0 

4)4 '0 M 0 (1) (/) 4/3 (--1 

0 • (.4 4-1 )3.) 4-4 CU a) ..Ct 

to El Z l. rd P4 .4 

13.1 

y, f:G 

W 

CI 

.--, 

'CI ,..-, .---. 

CO 0 a) 

0 4.4 r-.4 

4-1 0 4 

..--, F-4 r'd 

t-4 0 sin, a) 

r-l. $.4 

44 0 PO 0 I-4 

..., ly 40 03 

0 0 C., E.44 

czl 0 •.-0( 0, ,..." 

,-.4 a 0 

-. 0 a ••1 CA 

4, ,..-• 00a CO 

04 CO el rni 

t0 0 ...) ..--, ...4 CO 

0 4.4 0 )",(4 0 114 (), 

I-1 t14 0 a.) $-, a a 

.1 ,(0 r-4 CL. -, 0 W 

CO (1) 1-.4 0 0 4.../ 

v•cl 4-4 (1) I-) 0 4-) CO 

41 z ta al, Pt, o 

,.0 0 4-1 0 ....■ COZ 

brD (11( CO 144 g 4.4 41 r-0( Q.1 

"(i.. ._4 ..,'". 4 0 ,. , -, 

15 "0 4.1 •1Z) 4-1 RI 0 4-4 

1.1 (11 I-I 4-) (1-1 CS) M 

44 4M Cl 01 0 4-, C4-/ r'*4 

0 t113 4.4 4-44 4.1 0 4 

0 *4-1 c/3 a) an .‹4 4-1 <4 

0 

1.4 4,4 

Z 

00 

.1-) 

4 

tg) 

..-1 

a) 

/-4 

CO 

., 

4-4 

•)"4 

C 

I-4 

a) 

.‘,) 

to 
Cs) 

Mach Number 

Axial Force Coefficient, 7 seg-120, Stage 0 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

S D-CR-65-2 6 111-67 

CO 

C•1 (,1 e-1 o 

0 quoiamaop ealoa Twpcy 
Ax ial Force Coefficient, 2 CS-156, Stage 0 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-68 

SSD-CR-65-206 

00 O 

yD luaToTgjeoD 9010a reTxv 

r-i 1-1 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-69 Mach Number Axial Force Coe fficient, 3 CS-156, Stage 0 

cc 

1• F-1 

CO 0 

C•1 

'D 

,

luaToT,TgooD aDapa ieTxy 
4 7 8 

Mach Number 

9 10 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

111-70 

Fig. 111-49 Axial Force Coefficient, Stage I, All Vehicles 
my 

-206 1V-1 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

IV. CONTROL SYSTEMS 

This chapter discusses the studies conducted pertaining to 

the flight control system (FCS) stability, and thrust vector control (TVC) system sizing. The results of the vibration analysis 

that determined the mode shapes and frequencies of the flexible 

vehicle are presented. Basic problems that were evaluated arc: 

1) The effect of the large solid rocket motors (SRM) and 

large payloads upon the FCS; 

2) The requirements imposed on the TVC system in terms 

of total injectant required and maximum side force 

requirements. 

A. CONTROL SYSTEM STABILITY 

The primary objective of the flight control analysis was to 

determine if any limitations existed on the MOL payload size or 

the type of solid motors due to flight control stability. The 

problems studied to determine restrictions to the payload and/or 

the solid motor configuration were: 

1) Basic stability of Stages 0, I, and II; 

2) The load-relief capability of the Stage 0 FCS; 

3) The coupling of actuator and structural bending 

effects for Stages I and II. 

In most situations it is more desirable to restrict and limit 

the FCS than the payload size or type of solid motors. Therefore, 

the second objective of the study was to determine what restrictions and requirements must be applied to the FCS to control the 

various payload and solid motor configurations. 

The following discussion is subdivided into two main sub-

, sections. The first subsection explains all limitations and 

requirements imposed by the stability of the vehicle on the payload, SRMs, and FCS, 

iihwiSSWED 
The second subsection discusses in more detail the analysis that 

was performed and the problems encountered in the analysis. The 

lest two sections summarize the systems designed in this study 

and the criteria used in the analysis. 

1 Summarof S stem Re*uirements Im osed b ht Control 

Considerations 

The flight control studies showed that with respect to stability and load-relief capability of the FCS there is little difference between the three solid rocket configurations.. As the payloads and solids become heavier the structural bending mode frequencies become lower, and present a somewhat more difficult stability situation. However, as shown in Fig. IV-1, the structural 

frequencies for the three solid configurations are near each 

other; Therefore each configuration presents essentially the 

same stability problems. This point is further confirmed by 

observing the similarity of the autopilots designed for each 

of the solid configurations and their associated payloads (tabulation of the autopilot configurations is presented in Chap, IVIV.A.4). 

The studies have shown that there is very little difference 

between the maximum and the minimum length payload for any one 

of the three solid configurations regarding control system stability. However, as the payload becomes longer the stability 

problem tends to be slightly more complex. This is shown by the 

corresponding decrease in load-relief capability of the FCS 

(Fig. IV-2), The slopes of the lines in Fig. IV-2 are rather 

small, showing that there is a slight difference in the loadrelief capability between the longest and shortest payloads for 

any one solid configuration. 

The FCS studies can be extrapolated to longer payloads than 

those studied. This extrapolation is, however, limited to an 

additional 10 ft. The stability problems will not become significantly more complex for the 10 ft range of longer payloads 

than were observed in the present study. Therefore, it can be 

expected that these extrapolated payloads will possess essentially the same amount of load relief as given by extrapolating 

Fig, IV-2. The limitation exists because extrapolation beyond 

10 ft presents certain stability problems that the present analysis has not adequately investigated. These stability problems 

can. be solved, but it is questionable that the same degree of 

load relief can be maintained, 

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IV-2 
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IV-3 Yaw Structural Mode Frequencies; Flight Time = Liftoff, with Transtage 

44 

(oas/pea) Aouenbaaa epow lyinlnals 

Payload Length (ft) 
0 0 0 

(%) 42ob uT uoT1Dnpas 

0 

tt) 

Load Redu

c

tion bythe PCS,with Tran

stage 

N 

1-1 

bD 

or' 

0 N 

P".1 

GO 

r-, 

0 

Payload Leng

th (ft) 

0 

0 

0 

r-1 

Cr) V 

N 

0 

q> 

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IV-4 
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IV-5 

The FCS studies have defined an autopilot configuration for 

Stages 0, I, and II that will reasonably meet all stability criteria as defined by the specification requirement, This configuration is considered to be a minimum required capability autopilot 

since it is not known at this time what effect tolerances will 

have on the FCS. Tolerance problems may dictate additional capability and changes in the FCS. 

Stage 0 - The autopilot concept presently being used for the 

MOL/Titan III Stage 0 is similar to the one used in the Titan IIIC 

program. A block diagram of the pitch channel autopilot is presented in Fig. IV-3, The only major deviation from the present 

Titan IIIC autopilot is the requirement for a filter change when 

the load-relief system is removed at about 80 sec of flight time, 

Also there is the possibility of a requirement for one more additional gain change between 80 sec and Stage 0 burnout to optimize 

the system configuration, 

Due to the significant amount of load relief provided by the 

FCS, particularly in the yaw plane, the transient is significant 

when load relief is removed. Future studies may show that this 

transient must be eliminated, or reduced, through additional autopilot mechanization. 

Stage I - The autopilot for Stage I MOL at the pre,:ent is 

essentially the same as Stage I autopilot of the Titan- IIIC program. The only difference is that there can be no filter sharing 

with Stage 0 filters. Each stage will require independent filters. 

Stage II - In Stage II the first structural mode presents a 

significant phasing problem. The method selected to solve the 

problem was similar to the Titan IIIB approach, i.e,, insertion 

of two underdamped quadratic filters in the rate feedback loop, 

In addition, one in-flight gain change will be needed near Stage 

II midflight. 
IV-6 

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Fig. IV-3 Pitch Autopilot 

C4 

-4 

cb 

1 

rg4 

1 
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IV-7 

The sensor locations selected for the MOL configurations with 

transtage are: aft rate gyro, Station 887; forward rate gyro, 

Station 112; and accelerometer, Station 112. The reason for these 

sensor locations is demonstrated in Fig. IV-4. At Station 112 the 

amplitude of the first structural mode is small, allowing large 

accelerometer channel gains, and the slope of the second structural mode is small, thus minimizing second mode effects in the 

rate signal. For the configurations without transtage, the mode 

shapes are virtually unchanged, and sensor locations between Stations 235 and 257 are equally applicable, It is important to 

point out that these sensor locations are satisfactory for all 

three solid configurations and their associated payloads. For 

longer payloads than the ones studied these sensor locations will 

probably also be satisfactory. However, this depends on what 

shape the second mode takes when it reaches a frequency low enough 

to couple with the fluid slosh modes. For shorter payloads than 

those studied, the sensor locations will work well until the payload becomes less than 45 ft. 

The Stage II main engine actuator presently being used on the 

Titan IIIC program will not suffice for the MOL vehicle due to the 

structural bending and actuator coupling. The FCS studies show 

that this problem can be eliminated if the Stage II actuator and 

engine changes used on the Titan MB are used for the MOL booster. 

If the Titan IIIB actuator is used, the minimum 'Stage II engine 

frequency that can be allowed is 7,5 cps. 

For Stage I utilizing a 15:1 expansion ratio engine, the structural bending and actuator coupling is not critical. The minimum 

engine frequency that can be allowed is 8.0 cps. 

It is recommended that closed-loop guidance not be used during 

Stage 0, Closed-loop guidance could result in significant loss 

in load relief. This would be due to: 

1) Changes in autopilot gains to offset loss in aerodynamic margin due to guidance; 

2) The response of the vehicle to guidance signals in 

opposition to load relief. 

The most significant effect on the FCS of removing the transtage is that the forward-located sensors will have to be relocated 

farther back in Stage II. The effect of relocating sensors was 

investigated with the result that this relocation will have only 

a small effect on load-relief capability and basic stability of the 

vehicle, 
0 z

8 

0 a) 

0 q:) 

Vehicle Station(in.) 

Typical First 

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IV-8 

0 

7-4 

0 

0 

0 0 

co 

4 

1-4 

0 

0 ci 

0 

0.020immamill MIIIIIIIIIII 1 , ., . . : . . 

. 111 

, 

0.010 RIC-7. -erji , i' ' i Mill i . 

NIIIIIIIIIIbhb,. 1 ' 

, 

11 i i 

,, MIIIIIIIIMIlit: ' I.Indeformed Centerline .. 111111111111111 i ,-1 

0 IIIIEIIEINIIIIIIIIIEIIIMIIIIIIIIIIIMIMIIIII 

(887)* 

111111111111 11.1=111111 First Mode lill00111, 

—, 

'... .:. 1. * ' -1%.14.11.1.1111111.11111111 111 1.1 1.11M11 . 1 . • 

1111111111 0111111111111111111.1111111111111111111 ; 

.. . 

• 

- -c. -fY- 111111111111111111111111111111.1111111111111111111MINI ! . 

rd 11111 111 III Mill 

0030 

m Second Mode 1116 

11111111 111111 ill . 111 111 1111.111 -. . MII 441 

*Accelerometer and rate gyro location(with transtage) Mill111111W1111 

III 

1 

. milimibrft-.... -mod 11111 -0.040 1111111111111111111111111111111111111111111111111111 IIIII. 
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2, Flight Control Study Analysis 

Stage 0 - The approach taken in the design of the Stage 0 FCS 

was to maximize the load-relief capability of the control system, 

Load-relief maximization was performed by designing the FCS to the 

open-loop margin requirements only. These requirements are presented in Chap, IV.A.4, following. Further optimization of the 

autopilot design will not increase the load-relief capability, 

but will improve the stability margins to the design objective 

values. Consequently, more effort spent on optimizing the autopilot will not result in any significant amount of additional 

load relief. 

The stability problems encountered in Stage 0 were similar 

for all payloads and all three solid rocket configurations. 

Therefore, the following discussion of problem areas will apply 

to all payloads and solid configurations under study, 

One of the major problems associated with the large payloads 

and solids is the low frequencies of the structural bending modes, 

Figure IV-1 shows frequencies of the first and second modes as 

a function of payload length. To stabilize these low frequency 

modes, low frequency filtering is required, This is depicted by 

the autopilot filters tabulated in Chap. IV,A.3. This filtering 

results in significant phase lag in the rigid bOdy region, which 

intensifies the problem of maintaining the rigid body Margins. 

The low frequencies of the structural bending modes also produced considerable difficulty in maintaining the criteria of the 

closed-loop roots. The low frequency of the higher modes (second 

to the fourth) resulted in rather large bending amplitudes, To 

stabilize these modes to the 10-db requirement, the autopilot 

gains had to be kept as low as possible, The lowering of the 

gains made the closed-loop root criteria difficult to meet. In 

a few cases the maximum closed-loop root frequency that could 

be obtained was less than the criteria of 1.5 rad/sec, The lowest value obtained was 1,3 rad/sec, One reasonably successful 

scheme that was tried to solve this problem was a filter change 

at 80 sec (when load relief was removed). If closed-loop frequencies of 1.5 rad/sec or greater are required during portion of 

Stage 0 flight without load relief, then an in-flight filter 

change will definitely be required for Stage O. 
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To minimize the stability problems of the acceleration feedback it is best to locate the sensor near a mode of the first 

structural bending mode, or at a position of small first mode 

displacement. Also, because it allows for more filtering, it 

is desirable to locate the accelerometer so that the first mode 

bending is positive, In Fig, 111-4 the sensor locations selected for the MOL vehicles are shown. The forward rate gyro 

was located at Station 112 instead of Station 320 as in the 

Titan IIIC. This was to allow for a rate gyro for Stage III 

and to help solve the stability problems of Stage II. 

The reduction in loads obtained for each FCS configuration 

is presented in Fig. IV-2. The primary reason for the significant amount of load relief is the relatively large Kv. In previous MOL studies Kvwas kept low due to a low frequency instability that can occur when Kv is increased. The previous MOL 

studies used the ground rule that this low frequency root must 

be stable. For this study this ground rule was deleted because 

this root, even though unstable, has a time constant of more 

than 120 sec and does not cause any appreciable problem. With 

this ground rule discarded, Kvcan be increased significantly 

over what has previously been observed in MOL studies. 

The only significant problem caused by increasing'Kvis that 

the attitude error during load relief will be large. The trajectory studies show attitude error for the MOL vehicles with 

these Kv's to be around 5 deg at wind shear peak. Due to this 

large attitude error, the transient when load relief is removed 

at 80 sec is significant. If this transient becomes a problem 

in later studies, it can be solved by additional autopilot mechanization. This autopilot mechanization will be of a nature 

that will not affect the load-relief capability of the system. 

An example of this would be ramping out the gains when load relief is removed. 
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IV-11 

Another reason load relief of the FCS is large is because the 

first bending mode signal sensed by the accelerometer is small and 

easily filtered out. With no bending, the only limitations to 

the accelerometer gain are rigid body considerations which, in 

general, are not as constraining as structural bending. 

Typical stability plots for the 60-sec cases are included in 

Fig. IV-5 thru IV-16. Two general configurations were considered 

for MOL, with transtage and without transtage, The bulk of the 

flight control analysis was done on the configuration with the 

transtage. If the transtage is removed the primary effect on the 

FCS will be the relocation of the forward sensors. To determine 

this effect, a study was made with the forward-located rate gyro 

and accelerometer relocated at Station 235. The results of that 

study showed that there was no significant difference between the 

autopilots for the two sensor locations. The only difference was 

that the rate gain ratio had to be changed, but this has no effect 

on the load-relief capability of the FCS. Therefore, the stability 

problems and load-relief capability of the FCS for the two configurations (with and without transtage) are essentially the same. 

Figures IV-17 thru IV-20 illustrate the comparison of the cases 

with and without transtage, 
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1V-12 

Autopilot Parameters 

= 0.745 F = 30 

= 0.245 FRi= 20,20 1 

= 0.070 F

R2 = 20,20 

= 0.302E-3 FA= 3,3,3 

= 0,135E-3 T w 7.5 

-60--- _ 

-140 -100 -60 -20 0 20 60 100 140 180 220 260 

Phase (deg) 

Fig. IV-5 Frequency Response Plot for Stage 0; Pitch, T = 60 sec; 

7 seg-120; 54.5-ft, 28,000-lb Payload 
-60 -20 0 20 60 100 140 

Phase (deg) 

220 260 

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UNCLASSIFIED 

IV-13 

Fig, 1V-6 Frequency Response Plot for Stage 0; Yaw, T go 60 sec; 

7 seg-120; 54.5-ft, 28,000-lb Payload 
Autopilot Parameters 

30 

l0 

-2U- 

-4O- 

-50h20 60 100 

Phase (deg) 

1 

 -100 -60 180 220 2-'60 

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1V-14 

Fig. IV-7 Frequency Response Plot for Stage 0; Pitch, T 60 sec; 

7 seg-120; 75,0-ft, 28,000-lb Payload 
IV -15 

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Autopilot Parameters 

Phase (deg) 

Fig. IV-8 Frequency Response Plot for Stage 0; Yaw, T cm 60 sec; 

7 seg-120; 75.0-ft, 28,000-lb Payload 
IV-16 

KD3= 0.5 F

D u 10 

= 0.56 F

R1 a10,10 

KR1 

KR2 0.05 F

R2 " 10,10 

K

A u 0.3E-3 FA a 3,3,3 

= 0.133E-3 T a 7.5 

20 

10 

ro 10 

-20 

-30 

-40 

-50 

b0140 -100 -60 -20 0 20 60 100 140 180 220 260 

Phase (deg) 

FSSD-CR-65-206 

40 

30 

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Fig. IV-9 Frequency Response Plot for Stage 0; Pitch, T - 60 sec; 

2 CS-156; 50,0-ft, 35,000-16 Payload 
IV-17 

60 100 

Phase (deg) 

20 140 180 220 260 

SSD-CR-65-206 

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40 

30 

20 

10— 

Fig. IV-10 Frequency Response Plot fpr. Stage 0; Yaw, T = 60 sec; 

2 CS-156; 50.0-ft, 33,000,,8 Payload 
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IV-18 

Autopilot Parameters 

40 

- 0 

-140 -100 -60 -20 0 20 60 100 140 180 220 

Phase (deg) 

260 

Fig. IV-11 Frequency Response Plot for Stage 0; Pitch, T = 60 sec; 

2.CS-156; 65.0-ft, 33,000-lb Payload 
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IV-19 

Autopilot Parameters 

KuE. 0.25 F

D 

10 

= 0.25 F

R 

= 10,10 

KR1 

KR2 = 0,015 F 10,10 

KA0.45E-3 FA 

Ky 0.20E-3 

= )3)3 

I = 7.5 

4.82 

0,49 

1.2 

0 

40 

0 

0 

-20 

-40 

-50 

140 -100 -60 -20 20 60 

Phase (deg) 

140 Lib 220 260 

Fig, IV-12 Frequency Response Plot for Stage 0; Yaw, T = 60 sec; 

2 CS-156; 65.0-ft, 33,000-lb Payload 
-20 

-40 

6O t_.... 

-140 - 00 140 180 220 260 

1 

-20 0 20 60 

Phase (deg) 

00 -60 

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RELEASE 1 JULY 2015 

1V-20 

Autopilot Parameters 

= 0.67 

D 

= ,10 

KR1 

- 

0.33 

F

R1 = 15,15 

KR2 = 0.06 F

R2 = 15,15 

0.24E-3 F 3 5" 10 

A 

= 9.062E-3 T = 7.5 

0 0.60 

0 

ci 

,6 

40 

0 

Fig. IV-13 Frequency Response Plot for Stage 0; Pitch, T = 60 sec; 

3 CS-156; 50.0-ft, 42,000-lb Payload 
Autopilot Parameters 

KD= 0.28 E = 10 

KR1 13'28FR1 

 10,10 

KR2 = 0.28 Flu = 10,10 

K

A 

= 0.51E-3 

Kv= 0.23E-3 T 7.5 

40 

30 

20 

10 0.41 

5.73 

-20 0 20 60 100 140 180 220 —260 

Phase (deg) 

-30 

-50 

-60L-- - 1- 

-140 -100 -60 

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RELEASE 1 JULY 2015 

IV-21 

Fig. IV-14 Frequency Response Plot for Stage 0; Yaw, T = 60 sec; 

3 CS-156; 50.0-ft, 42,000-16 Payload 
40 

30 

20 

-40 

-50 t 

Autopilot Parameters 

KD= 0.48 

K

R1 

= 0.426 

K = 0.053 

R2 

K

A 

= 0,213E-3 

K

V = 0,035E-3 

F

D 

= 10 

F

R1 

= 20,20 

F

R2 = 10,10 

F

A 

 = 5

'

5

'

5 

T = 7,5 

'?140 -100 -60 -20 20 60 100 

Phase (deg) 

Fig, IV-15 Frequency Response Plot for Stage 0; Pitch, T = 60 sec; 

3 CS-156; 72,0-ft, 42,000-lb Payload 

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1V-22 

eimioa 

SSD-CR-65-2 
-30 

-40 

-50 

60 100 140 L80 220 260 

0 

-140 -100 -60 -20 20 

Autopilot Parameters 

40 

20 

0.42 

1.5 

3.8 

10i 

0 

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IV-23 

SSD-CR-65-206 

Phase (deg) 

Fig. IV-16 Frequency Response Plot for Stage 0; Yaw, T = 60 sec; 

3 CS-156; 72.0-ft, 42,000-16 Payload 
IV-24 

Autopilot Parameters 

5.45 

1 

40 

30 

20 

10 

0 10 

-20 

-30 

-40 

-50 

-60-- -I 

-140 -100 -60 -20 0 20 60 100 

Phase (deg) 

140 180 220 260 

SSD-CR-65-206 

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Fig. IV-17 Frequency Response Plot for Stage 0; Pitch, T img 80 sec; 65,0-ft, 

33,000-lb Payload; Sensors at Stations 235 and 887 
Note: 1, Before gain change. 

2, 2 CS-156. 

140 180 220 260 -60L- 

-140 -100 -60 

IV-25 

KD= 0.25 

K

R1 = 0.25 

KR2 0.015 

K

A 

= 0.45E-3 

Kv= 0.20E-3 

SSD-CR-65-206 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

Phase (deg) 

Fig, IV-18 Frequency Response Plot for Stage 0; Yaw, T = 80 sec; 65.0-ft, 

33,000-lb Payload; Sensors at Stations 235 and 887 
IV-2G 

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RELEASE 1 JULY 2015 

40 T 

5,45 

201- 

10h-- - 

-401— 

A 1 

-140 -100 -20 0 20 60 100 

Phase (deg) 

Autopilot Parameters 

Fig. IV-19 Frequency Response Plot for Stage 0; Pitch, T = 80 sec; 65,0-ft, 

33,000-lb Payload; Sensors at Stations 112 and 887 

IINC! A.SSITIE0 
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IV-27 

-6O .. 

-100 -60 -20 0 20 60 100 140 180 220 260 

Phase (deg) 

Autopilot Parameters 

KD = 0.25 

KR1 = 0.25 

KR2 = 0.015 

K

A = 0.45E-3 

K = 0.2E-3 

F

D = 

F

R1 = 

F

R2 

F

A

= 

T = 

10 

10,10 

10,10 

3,3,3 

7.5 

Note: 1. Before gain change. 

2. 2 CS-156. 

Fig. IV-20 Frequency Response Plot for Stage 0; Yaw, T = 80 sec; 

65.0-ft, 33,000-lb Payload; Sensors at Stations 112 

and 887 UNCLASSIFIED 
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IV-28 

Stage I - The criteria used for the FCS analysis of Stage I 

is presented in Subsection 4 of this section (control system requirements). As in Stage 0 the stability problems are generally 

the same for all payloads. Therefore the following discussion 

will be for all payload configurations. As in Stage 0, the structural bending modes have rather low frequencies. These low frequencies present a problem of coupling of the first bending mode 

with the fluid slosh modes. When this occurs the structural model 

gives rise to two or more modes that are similar to a first mode. 

With multiple first modes the stability problem is more complex. 

This problem occurs to some degree in Stage 0, but generally the 

first mode is below the fluid slosh modes. 

The payloads on the configurations studied for Stage I were: 

75 ft long, 28,000 lb, and 72 ft long, 42,000 lb. The autopilots 

designed for these configurations are presented in Chap. IV.A.3. 

The proposed Stage I engines for two of the MOL vehicle configurations differ from the corresponding Titan III engines by the 

nozzle expansion ratio (15:1). This heavier nozzle results in a 

considerable increase in the engine moment of inertia about the 

gimbal (approximately a factor of three). 

As a result of the proposed engines it was desirable to examine 

the structural bending-actuator coupling as related to stability 

analysis. The configurations selected for the survey were representative of the 28,000-, 33,000-, and 42,000-lb payloads, respectively. Each configuration examined had the greatest length 

vehicle for the payload involved, and thus, the most severe structural bending problems. The criteria used for the study were that 

if any structural damping, after coupling with the actuator dynamics, dropped to 0.004 or less, then actuator or engine redesign would be needed. Using a 9 cps engine, all coupled structural damping was above the damping ratio limit of 0.004. 

A survey was also conducted to determine the minimum engine 

frequency that could be tolerated with the above criteria. Results of that survey showed minimum acceptable engine frequency 

of 8.0 cps. Figures IV-21 thru IV-24 present a typical set of 

stability plots for Stage I. 
0 

0 

-20 

-30 

-40 

-50 

Autopilot Parameters 

= 1.08 F

D = 15 

Closed-Loop Frequency = 1.53 

Closed-Loop Damping = 0.403 

7. 

I 

1 , 

-60 -20 0 20 60 100 140 180 220 260 

:Phase (deg) 

0 '100 

40 

0 

20 

to 

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IV-29 

SSD-CR-65-206 

Fig. IV-21 Frequency Response Plot for Stage I; Pitch/Yaw, T = Start; 

28,000-lb, 75-ft Payload 
40 

30 

Closed-Loop Frequency = 2.74 

Closed-Loop Damping = 0.520 

-20 0 20 60 100 140 180 220 260 

Phase (deg) 

-60 

-50 

60140 -100 

Autopilot Parameters 

1(1) 1 = 1,08 F

D = 15 

KR, = 0.24 F

R1 = 35,10 

KR2 = 0.50 F

R2 = 35,10 

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IV-30 

SSD-CR-65-206 

Fig. IV-22 Frequency Response Plot for Stage I; Pitch/Yaw, 

T = Midflight (Start + 119 sec), Before Gain 

Change; 28,000-1b, 75-ft Payload 
1.60 

40 

AmIppilot Parameters 

KD = 0.55 FD = 15 

= 0.053 F

R1 = 35,10 

K

R2 = 0.34 F

R2 = 35,10 

30, 

20 

Closed-Loop Frequency = 1.51 

Closed-Loop Damping = 0.416 

10 

-10 

-20 

-30-- 

-50 

01 

-140 -100 -20 0 20 60 100 

Phase (deg) 

9.45 

-60 140 180 220 260 

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IV- 31 

Fig. IV-23 Frequency Response Plot for Stage I; Pitch/Yaw, T = Midnight 

(Start + 119 sec) After Gain Change; 28,000-lb, 75-ft Payload 
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IV-32 

Closed-Loop Frequency 0 3.01 

Closed-Loop Damping 0.626 

i 

-60 -20 0 20 60 100 

Phase (deg) 

Fig. IV-24 Frequency Response Plot for Stage I; Pitch/Yaw, 

T Burnout; 28,000-lb, 75-ft Payload 

140 180 220 260 
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SSD-CR-65-206 T V -3) 

Stage II - The basic stability investigation of Stage II uncovered only one significant problem. With the Titan IIIC-type 

autopilot, the first mode phase was about 180 deg. To eliminate 

this problem, two underdamped quadratic filters in the rate loop 

were used. This resulted in a more desirable phasing for the 

first mode. To determine if this filter was sensitive to tolerances, a modified tolerance study was performed. The results of 

that study showed there would be no significant problems with 

respect to tolerance associated with this filter technique. 

A preliminary study of coupling between the first structural 

bending mode and the Stage II engine was accomplished through 

digital simulation. Results of that study indicate that excessive 

coupling will occur if the present Titan IIIC Stage II actuator 

is used. An effective (and necessary) means of decoupling is 

now conceived for the Titan IIIB Stage II actuator. The Titan 

IIIB actuator differs from the Titan IIIC actuator in the followmanner: 

1) The effective actuator Lever arm is increased from 10 

to 14 in.; 

2) The actuator pressure feedback gain is reduced from 

0.00182 to 0.00075 in.V1O-sec. 

A preliminary tolerance analysis indicated that, while the 

Titan IIIC actuator is not acceptable for any MOL configuration at 

any reasonable natural engine frequency, the Titan MB actuator 

is acceptable for all MOL configurations studied, if the natural 

engine frequency is not lower than 7.5 cps. 

This actuator analysis and conclusions are only valid for payloads of equal or smaller length and weight than the ones studied. 

If a payload configuration of greater length and weight than the 

one studied is selected for MOL, it is possible that even the 

Titan IIIB actuator will not be adequate to solve the Stage II 

structural bending coupling problem. Typical Stage II stability 

plots are shown in Fig. IV-25 thru 27. 

3. Autopilot Parameters 

The slight differences between the various configurations is 

further demonstrated by the similarity of the autopilot configurations required. The configurations defined in this study are 

shown in Tables IV-1 (Stage 0), IV-2 (Stage I), and IV-3 (Stage II). 
Closed-Loop Frequency 

Closed-Loop Damping 

40 

30 

20 

10 

0 23.4 

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IV-34 

Fig. IV-25 Frequency Response Plot for Stage II; Pitch/Yaw, T Start; 

7 seg-120; 65.0-ft, 28,000-113 Payload 
140 180 220 260 

Autopilot Parameters 

= 1.45 

KR = 0,755 

FD =(i 3 O + 

40 

30 

20 

10 

4.2 

1.0 + 2(0,4)S + S2 

15 2 

(15) 

Closed-Loop Frequency 2.04 2.04 

Closed-Loop Damping 0.35 2, 

10 

-20 

-30 

-40 

-50 

1 

0 A 

-140 -100 -60 

Phase (deg) 

-20 0 20 60 00 

8.4 

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1V-35 

Fig. IV-26 Frequency Response Plot for Stage II; Pitch/Yaw, T = Midflight; 

7 seg-120; 65.0-ft, 28,000-16 Payload 
40 

30 

20 

-60--- 

-140 .100 

1_1 

-60 -20 0 20 60 100 140 180 220 24,( 

r 

Autopilot Parameters 

1 (15)2

)

2 

F 1.11 4. 2(0.

5 

4)S S

2 

Kt = 1,45 

K

R 

= 0.755 

F = (1.0 + A—) 

D 25/ 

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IV-36 

Phase (deg) 

Fig. IV-27 Frequency Response Plot for Stage II; Pitch/Yaw, T = Burnout; 

7 seg-120; 65.0-ft, 28,000-16 Payload 
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IV-37 

Table IV-1 Stage 0 Flight Control System Configurations 

Control 

System 

Configuration Constant 

Constants vs Flight Time 

0 to 35 sec 35 to 80 sec 

80 sec to 

Burnout 

7 seg-120; 54.5-ft, 28,000 -lb Payload 

Pitch Axis KD 0.505 0.745 0.311 

KR1 0.202 0.245 0.116 

K

R2 0.059 0.070 0.098 

K

A -- 0,302E-3 -- 

Kv 0.135E-3 -- 

FD 30 30 30 

F

R1 20,20 20,20 8,8 

F

R2 20,20 20,20 8,8 

F

A 3,3,3 -- 

-r 7.5 

Yaw Axis KD 0.80 0.29 0.55 

KR1 

0.32 0.26 0.23 

KR2 0.10 0.052 0.063 

KA 

.... 0.52E-3 ..- 

0.30E-3 -- 

FD 30 30 30 

F

R1 35,35 35,35 35,35 

F

R2 35,35 35,35 35,35 

7.5 
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IV-38 

Table IV-1 (cont), 

Control 

System 

Configuration Constant 

Constants vs Flight Time 

0 to 35 sec 35 to 80 sec 

80 sec to 

Burnout 

7 seg-120; 65.0-ft, 28,000-16 Payload 

Pitch Axis KD 0.45 0.63 0,35 

KR1 0.24 0.24 0.16 

K

R2 0.08 0.055 0.094 

K

A 

-- 0.19E-3 -- 

KV 

-- 0.09E-3 -- 

F

D 

20 20 20 

F

R1 15,15 15,15 10,10 

F

R2 15,15 15,15 10,10 

F

A 

- 5,5,5 -- 

T _.. 7.5 -- 

Yaw Axis KD 0.81 0.32 0.55 

K

R1 0.40 0.23 0.23 

KR2 0.07 0.037 0.063 

K

A 

0,53E-3 

Kv -- 0,28E-3 -- 

F

D 30 30 30 

F

R1 35,35 35,35 35,35 

F

R2 35,35 35,35 35,35 

F

A 

- 3,3,3 

T -- 7.5 - 
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IV-39 

Table IV-1 (cont) 

Configuration 

Control 

System 

Constant 

Constants vs Flight Time 

0 to 35 sec 35 to 80 sec 

80 sec to 

Burnout 

i 7 seg-120; 75.0-ft; 28,000-16 Payload 

Pitch Axis KD 0.40 0.66 0.39 

KR1 0.28 0.29 0.20 

KR2 0.10 0.052 0.092 

K

A - 0.212E-3 -- 

Kv -- 0.094E-3 

F

D 12 12 12 

FRI 10,10 10,10 10,10 

FR2 10,10 10,10 10,10 

FA -- 5,5,10 

T -- 7.5 -- 

Yaw Axis KD 0.82 0.35 0.55 

KRI 0.48 0.20 0.23 

KR2 0.042 0.022 0.063 

KA -- 0.54E-3 

KV 0.27E-3 .... 

F

D 30 30 30 

F

R1 35,35 35,35 35,35 

FR2 35,35 35,35 35,35 

F

A -- 3,3,3 

T -- 7.5 
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IV-40 

Table IV-1 (cont) 

Configuration 

Control 

System 

Constant 

Constants vs Flight Time 

0 to 32 sec 32 to 80 sec 

80 sec to 

Burnout 

2 CS-156; 50-ft, 33,000-lb Payload 

Pitch Axis KD 0.5 0.5 0.37 

KR1 0.56 0.2 - 

KR2 0.06 0.05 0.08 

K

A 

-- 0.32-E -- 

Kv 0.133E-3 -- 

F

D 

10 10 10 

F

R1 10,10 10,10 10,10 

F

R2 

10,10 10,10 10,10 

F

A 

-- 3,3,3 - 

7.5 -- 

Yaw Axis KD 0,504 0.25 0.5 

KR1 0.378 0.25 0.31 

K

R2 0.076 0.015 0.06 

K

A 

0.45E-3 

Kv 0.2E-3 -- 

F

D 

10 10 10 

F

R1 10,10 10,10 10,10 

F

R2 10,10 10,10 10,10 

F

A 

-- 3,3,3 -- 

7.5 
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IV-41 

Table IV-1 cont) 

Configuration 

Control 

System 

Constant 

Constants 

0 to 32 sec 

vs Flight Time 

32 to 80 sec 

80 sec to 

Burnout 

CS-I56; 58.5-ft, 33,000-lb Payload 

Pitch Axis KD 

KR1 

KR2 

K

A 

F

D 

F

R1 

F

R2 

F

A 

0.563 

0.427 

0.0675 

-- 

9 

15,15 

15,15 

0.517 

0.47 

0.0435 

0.195E-3 

0.0865E-3 

9 

15,15 

15,15 

3,3,3 

7.5 

0.423 

0.226 

0.08 

9 

15,15 

15,15 

Yaw Axis BL D 

KR1 

KR2 

K

A 

F

D 

F

R1 

F

R2 

F

A 

0.504 

0.378 

0.076 

10 

10,10 

10,10 

0.25 

0.25 

0.015 

0.45E-3 

0.2E-3 

10 

10,10 

10,10 

3,3,3 

0.5 

0.31 

0.06 

10 

10,10 

10,10 
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IV-42 

Table IV-1 (cont) 

Configuration 

Control 

System 

Constant 

Constants vs Flight Time 

to 32 sec 32 to 80 sec 

0 sec to 

Burnout 

CS-156; 65.0-ft, 33,000-lb Payload 

Pitch Axis 0.625 0.535 0.475 

1 

0.375 0.38 0.252 

l(R2 0.075 0,037 0.079 

K

A 

0.09E-3 

Kv 0.04E-3 

F

D 

8 8 8 

F

R1 

20,20 20,20 20,20 

F

R2 

20,20 20,20 20,20 

F

A 

3,3,3 

7.5 

Yaw Axis 0.504 0.25 0.5 

0.378 0.25 0.31 

0.076 0.015 0.06 

K

A 

0.45E-3 

0.2E-3 

F

D 

10 10 10 

F

R1 10,10 10,10 10,10 

F

R2 10,10 10,10 10,10 

F

A 

3,3,3 

7,5 
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IV-43 

Table IV-1 (cont) 

Configuration 

Control 

System 

Constant 

Constants vs Flight Time 

0 to 28 sec 28 to 80 sec 

80 sec to 

Burnout 

Pitch Axis 

3 CS-156; 50.0 

Kr) 

KR1 

KR2 

K

A 

KV 

F

D 

F

R1 

F

R2 

F

A 

-ft, 42,000 

0.52 

0.35 

0.075 

NV Mt 

oa 

10 

15,15 

15,15 

-lb Payload 

0.67 

0.33 

0.06 

0.24E-3 

0.062E-3 

10 

15,15 

15,15 

3,5,10 

7.5 

0.34 

0.18 

0.09 

10 

8,8 

8,8 

Yaw Axis 

KD 

KR1 

K

R2 

K

A 

F

D 

F

R1 

F

R2 

F

A 

0.64 

0.32 

0.095 

10 

10,10 

10,10 

0.2R 

0.28 

0.028 

0.51E-3 

0.23E-3 

10 

10,10 

10,10 

3,3,3 

7.5 

0.36 

0.28 

0.053 

10 

10,10 

10,10 
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IV-44 

Table IV-1 (cont) 

Configuration 

Control 

System 

Constant 

Constants vs Flight Time 

80 sec to 

0 to 28 sec 28 to 80 sec Burnout 

3 CS-156; 61.0-ft, 42,000-lb Payload 

Pitch Axis KD 

0.535 0.575 0.39 

KR1 

0.285 0.378 0.19 

KR2 

0.0875 0.0565 0.1 

K

A 

0.227E-3 -- 

-- 0.0485E-3 

F

D 

10 10 10 

F

R1 

17,17 17,17 14,14 

F

R2 

13,13 13,13 9,9 

F

A 

3,5,10 -- 

T -- 7.5 -- 

Yaw axis KD 

0.56 0.265 0.34 

Kill 

0.31 0.265 0.265 

KR2 0.0835 0.0265 0.050 

K -- 0.48E-3 -- 

Kv -- 0.215E-3 -- 

F

D 

10 10 10 

F

R1 

10,10 10,10 10,10 

F

R2 

10,10 10,10 10,10 

F

A 

-- 3,3,3 -- 

7.5 -- 
0 sec to 

Burnout 

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IV-45 

Table IV-1 (cont) 

Configuration 

Control 

System 

Constant 

3 'S-I56; 

Constants vs Flight Time 

0 to 28 sec 28 to 80 sec 

.0-ft, 42,00 -lb Payload 

itch Axis K

D 

0.55 0.48 0.44 

KRi 

0.22 0.426 0.20 

0.1 0.053 0.11 

KR2 

K

A 

0.213E-3 

0.035E-3 

10 10 10 

F 20,20 20,20 20,20 

R1 

F

R2 

10,10 10,10 10,10 

5,5,5 

7.5 

Yaw Axis K

D 

0.48 0.25 0.32 

0.3 0.25 0.25 

KR1 

KR2 

0.072 0.025 0.047 

K

A 

0.45E-3 

0.2E-3 

F

D 

10 10 10 

F

R1 10,10 10,10 10,10 

F

R2 

10,10 10,10 10,10 

F

A 

3,3,3 

7.5 
Configuration 

Control 

System 

Constant 

Constants vs Flight Time 

0 to 28 sec 28 to 80 sec 

80 sec to 

Burnout 

Al 1 Configuration 

KD 

0.36 0.36 0.18 

KR 

0.19 0.19 0.10 

30 30 30 

F

R 

30 30 30 

Roll Axis 

K

D 

Displacement gain 

KR1 

Rate gain for rate gyro located at Station 887 (sec) 

KR2 

Rate gain for rate gyro located at Station 112 (sec) 

K

A 

Accelerometer gain for the LASS located at Station 112 

(rad/in./sect) 

KV 

Velocity gain for the LASS located at Station 112 (rad/in./ 

sec) 

Displacement loop dynamics (rad/sec)* 

F

R1 Loop dynamics for KR1 (rad/sec)* 

F

R2 Loop dynamics for K_R2 (rad/sec)* 

F

A 

Accelerometer loop dynamics (rad/sec)* 

Integration constant (sec) 

*Given as corner fre uencies of first order filters 

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IV-46 

Table IV-1 (concl) 
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IV-47 

Table IV-2 Stage I Flight Control System Configurations 

Configuration 

Control 

System 

Constant 

Constant vs Flight Time 

Start to 

Midflight 

Midflight 

to 

Burnout 

75.0-ft, 28,000 lb Payload 

Pitch and Yaw Axes KD 1.08 0.055 

KR1 0.24 0.053 

KR2 0.050 0.34 

F

D 

15 15 

F

R1 10,35 10,35 

F

R2 10,35 10,35 

72.0-ft, 42,000 lb Payload 

Pitch and Yaw Axes KD 1.32 0.47 

KR1 0.33 0.05 

KR2 

1.10 0.33 

F

D 15 15 

F

R1 10,35 10,35 

F

R2 10,10 10,10 

Displacement gain 

1 Rate gain for rate gyro located at Station 887 (sec) 

2 Rate gain for rate gyro located at Station 112 (sec) 

Displacement Loop Dynamics (rad/sec)* 

F R1 Loop Dynamics for K (rad/sec)* R1 

F

R2 Loop Dynamics for KR2 (rad/sec)* 

*Given as corner frequencies of first order filters. 
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IV-48 

Table IV-3 Stage II Flight Control System Configurations 

Configuration 

Control 

System 

Constant 

Constant vs Flight Time, 

Start to Burnout* 

65.0-ft, 28,000-16 Payload 

Pitch and Yaw Axes 

KR2 

F

D 

F

R2 

1.45 

0.755 

(1.0 + 8/25) 

[1.0 + 2 (0.4) S/15 +S2 + (15) 

Configuration 

Control 

System 

Constant 

Constant vs Flight Time 

Start to 

Midflight 

Midflight 

to 

Burnout 

72,0-ft, 42,000-lb Payload 

Pitch and Yaw Axes KD 

KR2 

F

D 

F

R2 

1.8 

1.0 

(1.0 + S/25) 

1 

1,0 + 2(0.5)S/12 

1 2 

+S/(12)] 2 

_ 

1.4 

0.8 

(1.0 + S/25) 

[1.0 + 2(0.5)S/12 

0 2 

+S

2

//(12)1 

Kt) Displacement gain 

KR2 

Rate gain for rate gyro located at Station 112 (sec) 

K. Displacement loop dynamics (rad/sec) 

F

R2 Loop dynamics for KR2 (rad/sec) 

*No gain change was used in the 28,000-lb payload configuration 

studied, but margins were barely met. A gain change is recommended for a final autopilot design. 
Frequency Range 

Gain Margins 

Design 

Requirement 

Design 

Objective 

0 to 1st Structural Bending Mode 

(db) 

1st to 3rd Structural Bending 

Modes (db) 

Structural Modes above 2nd Mode 

(db) 

5.0 

8.0 

10,0* 

6.0 

10.0 

10.0* 

Phase Margins 

Frequency Range 

Design Design 

Objective Requirement 

0 to 1st Structural Bending Mode 

(db) 25 30 

Above 1st Structural Bending 

Mode (db) 45 60 

Closed Loop Requirements 

Frequency Range 

All Pitch/Yaw (rad/sec)t 

*Independent of phase contribution. 

tNot required during load relief operation. 

Design 

Objective 

Design 

Requirement 

0.4 to 2.0 0.3 to 1,5 

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11/-49 

4. Control System Requirements 

The following control system requirements were extracted from 

the Detail Specification for Standard Space Launch Vehicle 

Standard Core, SSS-TIII-D-SLV/01A010 
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IV-50 

B. TVC SYSTEM SIZING 

Dispersion studies were performed to estimate the requirements 

for the TVC system. TVC fluid requirements and equivalent thrust 

vector deflection angles were predicted from these studies. 

1. Fluid Utilization 

Pertinent vehicle and environmental parameters were varied 

through estimated 3o ranges. The effect of the parameters on TVC 

fluid required for control purposes was noted in the dispersion 

studies. The fluid required for the vehicle and environmental 

parameters was then combined to yield the total amount of fluid 

required. 

The same assumptions and ground rules listed in the rigid 

body airloads section are applicable to this section also. The 

results are summarized in Table IV-4. 

Table IV-4 Summary of Fluid Requirements 

Vehicle SRM 

Configuration 

TVC Fluid Required for 

Control Purposes (lb) TVC Fluid 

Engine 1 Engine 2 Required (lb) 

7 seg-120 4,508 6,377 12,754 

2 CS-156 8,623 10,214 20,228 

3 CS-156 10,042 11,300 22,660 

The right-hand column in Table IV-4 assumes that each tank is 

separate with no common manifolding between the two tanks. Note 

that Fig. 1V-28, IV-29, and IV-30 portray only the sum of Engines 

1 and 2 and not the total fluid required. The total fluid required is twice that required by the engine using the maximum', 

since the maximum engine cannot be predicted before flight. A 

breakdown of the totals are shown in Tables IV-5 thru IV-7. These 

show the contribution from each of the variables considered. The 

major contributing factor in the sizing of the system is the maximum wind profile. This can be seen by referring to Fig. IV-28 

thru IV-30, where the fluid required for the major variations is 

plotted versus time of flight. Therefore, to assure that the TVC 

system is of sufficient size for both WTR and ETR launches, the 

'most severe 624A maximum wind criteria were used. The longest 

payload length for each SRM configuration was selected for investigation. The parameters considered in the sizing study are: 
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IV-51 

1) Nominal trajectory - The nominal performing vehicle 

will require a determinable amount of TVC fluid for 

the side-force characteristics given in Fig. 4.1.-1 

thru 4.1.11 of TM 5141/31-65-16 (Rev 1), "Titan III/ 

MOL Data Book," dated 27 August 1965. 

2) Thrust vector misalignment - A mislignment in the 

direction of the thrust vector in both the pitch and 

yaw vehicle planes requires an additional amount of 

fluid. The variation is ±0.25 deg in both pitch and 

yaw. 

3) Thrust vector offset - This variation accounts for 

effects whereby the effective thrust vector gimbal 

point is varied ±2in. in the lateral direction and 

±20 in. in the longitudinal direction. 

4) SRM temperature differential - The effects of a maximum temperature differential between the two SRMs of 

10°F were considered. 

5) SRM thrust deviation - The perturbational effects of 

considering maximum and minimum thrust-time histories based on the ±3a values supplied by the SRM manufacturer. 

6) SRM tailoff thrust differential - The differential 

thrust magnitudes during tailoff were considered as 

separate variations. These accounted for a considerable amount of fluid. 

7) Maximum wind - The 624A criteria were used for this 

variation. This is more severe than the WTR criteria 

and is the largest contributor to the fluid requirement. The wind azimuth used was 240 deg. 

8) Mean wind - The mean wind used was the 50% NASA wind. 

9) Roll maneuver - A maximum roll maneuver of 50 deg 

was considered. 
Maximum Total TVC Used 

Maximum Wind 

Pitch or Yaw TV isalignment 

NASA Mean Wind 

Temperature Differential and 

ust Dispersion 

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IV-52 

Time c ) 

Fig. IV-28 TVC Fluid Utilization, 7 seg-120 
20 it 

SSD-CR-65-206 

Note: . WTR launch. 

2. 80-n-mi circular orbit 

18 

16 

14 

12 

10 

6 

4 

TVC Fluid Used (103 16) 

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20 40 60 80 100 120 140 160 

Time (sec) 

Fig. IV-29 TVC Fluid Utilization, 2 CS-156 

IV-53 
0 20 40 60 80 100 

Time (sec) 

Fig. IV-30 TVC Fluid Utilization, 3 CS-156 

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1V-54 

22 I 1 

used 

20 

18 

16 

14 

-0.W) 

120 140 1CO 
Table IV-5 Injectant Requirements, 7 seg-120 

I. Independent Terms Incremental. lniectant 

Total A. Pitch Plane Engine 1 Engine 2 

Thrust Vector Misalignment 76 2,054 2,130 

Thrust Vector Lateral Offset 714 700 1,414 

Thrust Vector Longitudinal Offset 672 673 1,345 

. q p= Total Pitch Plane Required 

-IIEReqp 983 2,272 3,255 

B. Yaw Plane 

Thrust Vector Misalignment 1,060 1,061 2,121 

Thrust Vector Lateral Offset 714 700 1,414 

Thrust Vector Longitudinal Offset 672 673 1,343 

q = Total Yaw Plane Required 

- Y 

=2 Re4 1,444 1,438 2,882 

Q = Total Pitch and Yaw Required 

PY 

1( 2 2y 1,235 1,901 3,136 

cIP qY 

C. Maximum Wind-Mean Wind 2,004 2,852 4,856 

D. SRM Maximum Temperature Di erentlal 748 1,063 1,811 

E. SRM Thrust Dispersions 748 1,063 1,811 

F. SRM Tailoff Differential 1,106 1,148 2,254 

= Total Requirement of Independent Terms 

2 2 2 

PY + C +1) + 8' + F' 2,706 3,768 6,474 

II, Dependent Terms 

AA. Azimuth Orientation (50 deg) 416 224 640 

BB. Mean Wind (50% NASA) 670 1,533 2,203 

CC, Nominal Traiectory 716 852 1,568 

- Total Requirement of Dependent Terms 

= AA + BB + CC 1,802 2,609 4,411 

T = Total Injectant Requirements 

= T 4,508 

I 

+ T

II 

6,377 10,885 

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IV- 55 
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IV-56 

Table IV-6 TVC Iniectant Requirements, 2 CS-156 

I. Independent Terms Incremental Injectant 

A. Pitch Plane Engine 1 Engine 2 Total 

Thrust Vector Misalignment 88 3,110 3,198 

Thrust Vector Lateral Offset 1,179 1,201 2,380 

Thrust Vector Longitudinal Offset 1,210 1,201 2,411 

qp= Total Pitch Plane Required 

1,692 3,544 5,236 

B. Yew Plane 

Thrust Vector Misalignment 1,636 1,670 3,306 

Thrust Vector Lateral Offset 1,179 1,201 2,380 

Thrust Vector Longitudinal Offset 1,210 1,201 2,411 

q y= Total Yaw Plane Required 

.--..viEReq 2y 2,352 2,382 4,734. 

0 = Total Pitch and Yaw Required 

2 2

- 

' 

Op , 

2,049 3,019 5,068 

C. Maximum Wind-Mean Wind 3,228 3,876 7,1.04 

D, SRM Maximum Temperature Differential 

3,343 3,778 7,121 

E. SRM Thrust Dispersions 

F. SRM Tailoff Differential 2,659 2,683 5,343 

T = Total Requirement of Independent Terms 

2 + 2 5,733 6,753 12,486 

II. Dependent Terms 

AA. Azimuth Orientation (50 deg) 620 425 1,045 

BB. Mean Wind (50% NASA) 1,098 1,741 2,839 

CC, Nominal Trajectory 1,172 1,295 2,467 

= Total Requirement of Dependent Terms 

AA + BB + CC 2,390 3,461 6,351 

T = Total Injectant Requirements 

T = T 8,623 

I 

T

II 

10,214 18,837 
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IV-57 

Table IV-7 Injectant Requirements, 3 CS-156 

Total 

3,102 

2,763 

2,508 

5,36? 

4,041 

2,763 

2,508 

5,501 

5,506 

7,379 

8,707 

3,993 

13,322 

1,301 

2,019 

4,700 

3,020 

21,342 

I. Independent Terms Incremental Injectant 

Engine 1 Engine 2 

121 2,981 

1,358 1,405 

1,278 1,230 

1,869 1,516 

2,001 2,040 

1,358 1,405 

1,278 1,230 

2,735 2,766 

2,342 3,164 

3,092 4,287 

4,306 4,401 

1,993 2,000 

6,128 7,194 

810 491 

538 1,481 

2,566 2,134 

3,91.4 4,106 

10,042 11,300 

A. Pitch Plane 

Thrust Vector Misalignment 

Thrust Vector Lateral Offset 

Thrust Vector Longitudinal Offset 

q p = Total Pitch Plane Required 

=iftReq2 

B. Yaw Plane 

Thrust Vector Misalignment 

Thrust Vector Lateral Offset 

Thrust Vector Longitudinal Offset 

q= Total Yaw Plane Required 

=/2211cq2 

Q = Total Pitch and Yaw Required 

PY 

_L I 2 21 1` 

\qp qy ) 

2 

C. Maximum Wind-Mean Wind 

D. SRM Maximum Temperature Differential 

E , SRM Thrust Dispersions 

Terms 

Terms 

F. SRM Ti1off Differential 

T

I 

= Total Requirement of Independent 

=[

O2 +C2  +D

2 

 +E

2 

 +F 

PY 

II. Dependent Terms 

AA. Azimuth Orientation (50 deg) 

BE. Mean Wind (50% NASA) 

CC. Nominal Trajectory 

T

II 

= Total Requirement of Dependent 

= AA + BB + CC 

T

T 

= Total Inlectant Requirements 

= T

I 

+ T

II 

UNCLASSIFIED 
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The characteristics of the TVC system are included in the MOIL 

Data Book. These data include (1) side flow rate per quadrant vs 

command voltage with zero voltage in adjacent quadrants; (2) side 

force per quadrant vs command voltage for various motor vacuum 

nozzle centerline thrust and aft end stagnation chamber pressure; 

(3) nozzle centerline axial augmentation force per quadrant vs 

command voltage for various vacuum nozzle centerline thrust and 

aft end stagnation chamber pressures, and (4) nozzle centerline 

axial augmentation force during dump operation vs command voltage 

for various motor vacuum nozzle centerline thrust and aft end 

stagnation pressures. 

The amount of TVC fluid required for the variation in the above 

parameters was combined according to the following formula; 

TVC Fluid Required I1/2(RSS2

pitch+ RSS

2

Yaw 

+ (Maximum Wind - Mean Wind)2 + (SRM Temperature Differential)2 

+ (SRM Thrust Dispersion)2 + (SRM Tailoff)2 

] 

+ (Azimuth Orientation) + (Mean Wind) 

+ (Nominal Trajectory). 

The TVC fluid required by Engine 2 exceeded that of Engine 1 for 

all configurations studied. This is due primarily to the tendency 

of the vehicle to roll due to the rolling moment that results from 

the asymmetric TVC tank location. The cancellation of this rolling moment requires a differential deflection between Engines I 

and 2 in the pitch plane. 

Figures IV-28 thru 1V-30 show graphically the usage of TVC 

fluid for the various parameter dispersions. The top line curve 

labeled "Maximum Total TVC Used" represents the estimated 3a value 

of fluid required during flight, but not the fluid that must be 

loaded. Since the fluid required in each tank depends on wind 

direction, each tank must be loaded to the maximum expected amount, 

as shown in Table IV-4. 
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IV-59 

2. Side-Force Requirements 

The TVC side-force requirements were estimated by evaluating 

the six-degree-of-freedom trajectories used for the loads studies 

(including the wind search runs) and the six-degree-of-freedom 

trajectories used in the TVC utilization studies. 

The results of this evaluation showed that all SRM configurations can be controlled with the following equivalent thrust vector 

deflection capability: 

Equivalent Thrust 

Vector Deflection 

Flight Phase (deg) 

Liftoff to Clearing Launch 

Stand 3 

Clearing Launch Stand to End 

of Web Action 

During Thrust Tailoff 6 

C. VIBRATION ANALYSIS 

During this study, vibration analyses were conducted to provide the mathematical description of the flexible airframe for 

input to the FCS analysis and to the vehicle loads analyses. 

The most significant factor resulting from the vibration 

analysis that affected the FCS was the great reduction (compared 

to Titan IIIC with short payloads) of the structural mode frequencies. For example, Fig. IV-1 shows that there is almost a 

50% reduction in the first structural mode frequency compared to 

Titan IIIC (similar reduction in the higher structural mode frequencies also occur). This imposes problems on the FCS as discussed in Chapter IV.A. 

Typical Stage 0 first and second structural mode shapes are 

shown in Fig. IV-4. 
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V. STAGING 

This chapter presents a staging analysis of the Manned Orbiting Laboratory (MOL) performance improvement configuration described in Chapter II. The analySis is limited to solid rocket 

motor (SRM) separation. Titan IIIC staging evaluations have shown 

upper-stage separations to be relatively insensitive to upper 

body weight and, therefore, are not expected to create any unusual staging problems. The difference in Stage I engine thrust 

tailoff characteristics between a 15:1 engine compared to that of 

an 8:1 engine would have a small effect on the fire-in-the-hole 

staging technique. A staging analysis for each of these upper 

stages was deemed unnecessary for the purpose of this study and 

would serve little useful purpose before the MOL vehicle configuration is selected. Staging of the three SRM strap-on configurations, however, produces significant interaction with the MOL 

booster configuration selection. 

The three basic SRM strap-on configurations, i.e., 7 segment, 

120-in. (7 seg-120), 2 center segment, 156-in. (2 CS-156), and 3 

center segment, 156-in. (3 CS-156), are evaluated as to feasibility of separating the burned out SRMs from the Titan III core 

using the lateral translation technique successfully employed on 

Titan IIIC. This technique uses short-burn solid rocket motors 

mounted perpendicular to the SRM longitudinal axis to translate 

the SRM laterally away from Stage I. The staging rockets are 

located in the SRM nose fairing and tail skirt areas with a proper 

balance of orientation and thrust levels to obtain satisfactory 

clearance at separation. 

A staging sequence is selected that provides adequate control 

of the vehicle and respects structural limitations during SRM 

thrust decay. 

A. CONFIGURATIONS EVALUATED 

Of the many combinations of core configuration, payload, and 

SRMs possible, only the three shown in Table V-1, which represent 

significant staging problems, are evaluated. 
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V-2 

Table V-1 Configurations Evaluated 

SRM 

Option 

Stage I 

Engine 

Expansion 

Ratio Transtage 

Payload 

Weight (lb)/ 

Length (ft) 

Aerodynamic 

Pressure 

(psf) 

(qa) 

Product 

7 seg-120 

2 CS-156 

3 CS-156 

15:1 

8:1 

8:1 

Yes 

Yes 

Yes 

33,000/58 

33,000/58 

33,000/58 

40 

20 

5 

80 

40 

10 

Note that certain vehicle characteristics were assumed common 

to the three configurations so that the comparison of results 

might have a common basis. This is true of the payload weight 

(33,000 lb), payload length (58 ft), and total core weight 

(407,000 lb) at Stage I start. The thrust vector control (TVC) 

fluid injection system is assumed in each case to have the same 

side force to axial thrust relationship for a given thrust vector 

deflection command as the Titan IIIC system. 

The Titan IIIC staging rockets are used in the evaluation 

study. The nominal staging rocket thrust is 4740 lb, with a burn 

time of 2,6 sec. The rockets are stacked one over the other with 

their thrust vector oriented as nearly as possible through the 

projection of the SRM center of gravity (cg) in the roll plane. 

The staging rockets are canted approximately 30 deg outboard of 

the yaw plane to avoid plume impingement on Stage I. 

The staging sequence is predicated on the use of an inertial 

guidance system. The inertial measurement unit (IMU) senses a 

drop in longitudinal acceleration to a predetermined level after 

start of SRM thrust tailoff. This g-level sensing initiates the 

staging sequence with discrete signals being issued by the guidance 

system for Stage I ignition, Stage 0/I attachment ordnance, and 

staging rocket ignition. 

Table V-2 presents the mass property data used in the staging 

analysis. Aerodynamic coefficients used in the analysis include 

Stage I engine plume impingement forces and are documented in 

TM 0494/10-562. 

*Stavin.Aerod namics for MOL Com•atibili with 7-Se ment 

120-in. Diameter SRMs and 2- and 3-Segment, 156-in. Diameter SRMs. 

TM 0494/10-562. Martin Company, Denver, Colorado, 5 August 1965. 
ItI-CR-65-206 

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Table V-2 Staging Mass Property Data 

Step 0 Burnout Per Solid 

7 Segment 2 Segment 3 Segment 

Total Burnout Weight (lb) 94,629 122,780 148,090 

TVC Propellant at Burnout 

(lb) 4,250 6,000 6,555 

Center of Gravity (in.) 

Horizontal V.S. 866. V.S. 985. V.S. 895. 

Vertical W.L. 70.7 W.L. 70.2 W.L. 69.3 

Lateral B.L. 136.5 B.L. 134.4 B.L. 135.3 

Moment of Inertia (slug-ft2) 

Pitch 2.6 x 106 7.1 x 106* 4.6 x 1061 

Yaw 2.6 x 106 7.1 x 106 4.6 x 106 

Roll 0.07 x 106 0.12 x 106 0.15 x 106 

*Moment of inertia for 2-seg SRM was based on SRM manufacturer data. 

2-seg data appear questionable. 

t3-seg SRM moment of inertia data are based on Martin-generated data. 

Stage I Start Weight (33,000-16 Payload) 

Weight (lb) 407,000 

Center of Gravity (in.) 

Horizontal V.S. 645. 

Vertical W.L. 60.1 

Lateral B.L. 0.4 

Moment of Inertia 

(slug-ft2) 

Pitch 15.17 x 106 

Yaw 15.17 x 106 

Roll 0.034 x 106 

11-3 
B. STAGING CRITERIA 

The staging sequence and SRM separation are considered satisfactory if the following criteria are met: 

1) SRM separation will be accomplished after web burnout 

and adequate Stage I thrust has been achieved. For 

optimum performance, the SRMs must be jettisoned as 

soon as possible after their thrust-to-weight (T/W) 

ratio falls below the T/W ratio of Stage I. The extent of lateral translation staging rocket thrust 

must be sufficient to produce satisfactory separation 

clearance in the presence of one staging rocket failure either forward or aft, maximum SRM residual 

thrust, core plume force impingement, and adverse 

aerodynamic loads no greater than produced by the 

following conditions 

a) 7 seg-120 

Dynamic pressure, q = 40 psf 

Angle of attack, a = 2 deg 

Product, qa = 80 lb-deg/ft2 

b) 2 CS-156 

Dynamic pressure, q = 20 psf 

Angle of attack, a = 2 deg 

Product, qa = 40 lb-deg/ft2 

3 CS-156 

Dynamic pressure, q = 5 psf 

Angle of attack, a = 2 deg 

Product, qa = 10 lb-deg/ft2 

2) Axial acceleration must be maintained on the core 

vehicle during SRM thrust decay consistent with the 

minimum acceleration requirements of Stage I liquid 

engine net positive suction head (NPSH). 

3) Adequate control authority must be maintained to cope 

with maximum SRM differential thrust during the tail- 

• off period. 

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V-4 
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V-5 

4) Axial acceleration at Stage I ignition in the presence 

of maximum thrust differential must not violate Stage 

I structural limitations. 

C. STAGING SEQUENCE CONSIDERATIONS 

During the SRM thrust tailoff period, differential thrust 

between the two SRMs results in decreasing control authority. 

The Stage I engines should be ignited as soon as possible to augment control authority. The Stage I engine thrust buildup must 

not cause the booster acceleration to exceed the structural design limits of the core. The structural load limit thus determines the earliest time for Stage I ignition (87FS1). The EMU 

g-sensing level is established from the earliest time of core 

ignition after allowing for variations in SRM tailoff characteristics and guidance delays in sensing and issuing discretes. 

It is desirable for improved payload capability to jettison 

the SRMs as soon as their T/W ratio falls below the T/W ratio of 

the core. However, the effect of the canted residual SRM thrust 

vector is to rotate the aft end of the SRM into the core if 

residual thrust is too high at separation. The expected SRM 

thrust decay curves , therefore, determine the time of physical 

separation. 

Several aspects of the staging sequence are independent of 

the SRM configuration as indicated in the following typical 

staging sequence: 

1) Zero time reference - Sensing longitudinal acceleration initiates the staging sequence; 

2) 1.15 to 2.65 sec - Time from g-sensing to 87FS1 

(this represents the time delays in the airborne 

computer from sensing the prescribed g-level to 

issuance of a discrete); 

3) 0.7 to 1.1 sec - Time between 87FS1 and actual thrust 

rise; 

4) The time from 87FS1 

 to SRM separation is a function 

of the maximum residual SRM thrust and must be in 

multiples of the EMU computation cycle of 1.25 sec 

(assumed for this study). 
Location 

Forward/Aft 

Staging Rocket Requirements* (per SRM 

Number 

of 

Rockets 

4/4 

5/4 

4 

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V-6 

A typical profile of axial acceleration (T - A)/W vs time 

during the staging sequence is shown in Fig. V-1. The actual 

staging sequences selected for the three SRM configurations are 

also presented in this figure. Problems associated with these 

sequences are discussed in succeeding paragraphs. 

Typical Sequence 

87FS1 

r t 

1 

I 

1 

I 

Sensed 

1 1 

1 1 

Core 

„,,,,,,___f ___ 

Thrusting Acceleration 

Level , I 

SRM 

Separation 

Time-10- 

Staging Sequence 

SRM 

Option 

Sensed 

Acceleration' Time of 87FS1t 

(g) (sec) 

Time for SRMs 

Separation 

(sec)t 

7 seg-12O 1.90 1,15 --6-2,65 12.40 —10,- 14.00 

2 CS-156 2.29 1,15 —4.2.65 11.15'"' 12,75 

3 CS-156 2,29 1.15--s.2.65 8.65 —IP. 10.25 

*Titan IIIC lateral separation rockets. 

tTimes sequenced from sensed acceleration level major ADC 

comp, cycle, 1.25 sec. 

Fig. V-1 Staging Sequence 
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V-7 

D. STAGING SEQUENCE ACCELERATION PROFILE 

Acceleration profiles for the three basic SRM configurations 

are presented in TM 5141/31-65-19.* Each of these profiles is 

very similar to the typical profile shown in Fig. V-1. The two 

points of concern with any sequence involve the maximum structural acceleration limit at the time of thrust rise after 87FS1, 

as indicated by a rise in engine chamber pressure (Pc) and the 

minimum acceleration at SRM separation for NPSH requirements. 

These acceleration levels have been tabulated for convenience in 

Table V-3. 

Table V-3 Staging Sequence and Accelerations 

Configuration 

Sensed 

Acceleration (g) 

87FS1 P, Ri;t, SRM Ejoccion 

Time from 

g SensIng (sec) 

Ilme From 

87FS (sec 1 

Time 

' Level 

From 

g Sensing (sac) 

Minimum 8 and Stage I 

Propellant Temperature 

1.15 Min 0.7 Min 1.1 Min 12.4 Min • 0.860 70°F 

7 seg-I20 1.90 

2.65 Max Li Mx 1.4 Max 14.3 Max 0.850 9O* " F

.15 Min 11.7 Min 0.94 Min ILlS Min 0681 70° F 

2 CS-156 2.29 

2.65 Max 1,1 Max 1.85 Max 12.75 Max 0.676*: 90*P 

1.15 Min 0,7 Min I.13 Min 8.95 Mtn 0,625" 70'F 

9 CS-I56 2.29 

2.65 Max 1,1 Max 1.85 :lax 10,25 Max 0.620 90° F 

*Accelerations insufficlene to meet. present NPSH oxi dizer requirements of 44 ft. 

tAccelerationa insuffielent to meet proposed oxidizer requirement e of 3) ft, Lockup pressure moat be increased by 7 psi, 

*MOL Com.atibilit Solid Rocket Motor Sta n Anal sis Re ort. 

TM 5141/31-65-19. Martin Company, Denver, Colorado, August 965. 
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V-8 

1. Staging Loads 

The maximum acceleration levels at P

c 

rise combined with maximum SRM thrust differential shown in Fig. V-2 for an early Stage I 

engine start sequence produce significant structural loads in the 

core structure. The primary area of the core structure affected 

by staging loads is the Stage I longeron. The maximum longeron 

load actually occurs just after Stage I engine start where start 

transient maximum thrust overshoot occurs. The ultimate longeron 

loads at vehicle station 1225 for each of the SRM configurations 

are plotted in Fig. V-3 as functions of longitudinal acceleration 

and SRM differential thrust. Note that differential thrust vs 

longitudinal acceleration is shown for reference purposes. 

The maximum vehicle longitudinal acceleration at Stage I 

engine start is 1.6 g for the 7 seg-120 SRM and 1.85 g for both 

the 2 CS-156 and 3 CS-156 SRMs. These design conditions are taken 

from Table V-3. The longeron load for the 7 seg-120 configuration 

is approximately equal to the Titan IIIC test load. However, the 

loads for the 2 CS-156 and 3 CS-156 SRMs are approximately 207. 

higher than the test load and necessitate a redesign of the longeron as described in Chapter II. 

2. NPSH Requirements 

Referring again to Table V-3, an asterisk identifies the SRM 

configurations that do not meet the present oxidizer NPSH requirement of 44 ft at SRM separation. None of the three SRM configurations meet the present NPSH requirement when 90°F Stage I propellant temperatures are considered. The 90°F propellant temperature 

is compatible with Eastern Test Range (ETR) launch requirements. 

The maximum propellant temperature dispersion expected at the 

Western Test Range (WTR) is 75°F. The 7 seg-120 SRM and the 2 

CS-156 SRM both meet the present 70°F NPSH requirement. Reduction 

of the present oxidizer NPSH requirement of 44 ft to 35 ft has 

been proposed to the Aerojet-General Corporation. This proposal 

has been tentatively accepted and eliminates the need for any pressurization system redesign except for the 3 CS-156 SRM configuration using 90°F propellant temperature. An increase in the pressurization system lockup pressure of 7 psi is required for that 

configuration in addition to the proposed reduction in NPSH requirement. Further detail on the NPSH problem is presented in 

Chapter VII. 
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SSD-CR-65-206 

U 

a) 

us 

.■ 

a) 

S 

•■-t 

H 

(gT co") Isnot'', 
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V-10 

SSD-CR-65-206 

Legend: 

C5-156 

Factor of safety 

Longeron Load 

Thrust Differential 

Proposed Design Condition 600 

Note: 

-0 550 

540 

0 

oo 

4.) 

t--1 

500 ce) o 

o 

0 

V 11 

0 

0 .0 

)-4 

450 g § 

w " 

400 

350 

—207 > Test 

520 

500 

Limit Thrust D ifferential Between Solids 

p51 m 480 m Wes. 

4.,3 V 0 

1 121 

0 b 

4-10 0 

r..1 N4 41 

•—• 

460 

440 

1.2 14 1.6 1.8 20 

Limit Longitudinal g at Stage I Engine Start 

300 420 

Fig. V-3 Stage I Ultimate Longeron Load (Vehicle Station 1224) 

at Stage I Engine Start 
V-11 

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5;SD-CR-65- 06 

E. DIFFERENTIAL SRM THRUST TAILOFF 

The SRM tailoff curves shown in Fig. V-2 produce several 

important effects depending primarily on the time required to 

reach zero thrust and on the differential thrust between them. 

The effect of differential thrust on structural loads has already 

been discussed in Section D of this chapter. Differential thrust 

also results in decreased control authority and necessitates 

Stage I ignition to augment control available from the SRM TVC 

system. In this study, a TVC system is assumed in each case to 

have side force characteristics within the Titan IIIC state of 

the art. Figure V-4 shows vehicle yaw attitude response during 

SRM tailoff to be controllable and not unlike what is presently 

expected on Titan IIIC. 

The time required for SRM thrust to decay to essentially 

zero governs the time at which separation can occur. The 3 

CS-156 SRM has the steepest thrust decay slope and earliest time 

to zero thrust. The staging sequence for the 3 CS-156 SRN is 

therefore the shortest from g-sensing to separation. 

Attitude (deg) 

r 

Legend! 

 5 seg-120 

7 seg-120 

..... 2 CS-156 

— 3 CS-156 

Core Thrust Rise 

SRM s Staged 

8 10 12 14 16 18 20 

Time from End of SRM Web Action (sec) 

Fig. V-4 Control Authority during SRM Tailoff (Minimum/Maximum SAM Tailoff, 

Late Core Start) 
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V-12 

F. STAGE 0/I SEPARATION 

To demonstrate satisfactory clearance during SRM separation, 

different combinations of staging rockets are considered until 

the proper separation motion is achieved.. Plots of relative motion in the staging plane, based on the conditions and configurations shown in Table V-4, are presented in Fig. V-5 thru V-7. 

Table V-4 Summary of Separation Plots 

Configuration 

, 

SRM 

Tailoff 

Staging Rockets 

Forward Aft Sequence 

_ 

Figure 

7 seg-120 

Nominal 

Nominal 

Nominal 

Nominal 

Minimum 

Maximum 

5 

3 

4 

4 

4 

4 

5 

4 

5 

3 

3 

Early 

Early 

Late 

Late 

Early 

Early 

V-5(b) 

V-5(a) 

V-5(d) 

V-5(c) 

V-5(f) 

V-5(e) 

2 CS-I56 

Minimum 

Maximum 

Nominal 

Nominal 

Minimum 

Maximum 

5* 

5 

4 

5 

5 

5 

4 

4 

5 

3 

3 

Early 

Early 

Early 

Early 

Early 

Early 

V-6(b) 

V-6(a) 

V-6(d) 

V-6(c) 

V-6(f) 

V-6(e) 

, 

, 

3 CS-156 

Minimum 

Maximum 

Nominal 

Nominal 

Minimum 

Maximum 

5 

4 

5 

5 

5 

4 

4 

4 

5 

3 

3 

Early 

Early 

Early 

Early 

Early 

Early 

V-7(b) 

V-7(a) 

V-7(d) 

V-7(c) 

V-7(f) 

V-7(e) 

*Recommended number of staging rockets for satisfactory separation. 
7 seg-120 SRM Staging 

azi 00 

a) 

lzj 

r.4 

41) • 

'CS C•4 0 I 

N 

o cu

a) a 

0 

N 

1.1 

00 

a) 

N 

a) 

CC 

006 

II 

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V-13 

SSD-CR-65-206 

.C) inal-Nominal SRM Tailoff 

vs 

0 

00

CC

CC 

00 

a) 

Lri 

N 

otf 

T.) 

3 

0 1. 

4.1, 

C.% 

0 

00 

60 

n3 
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V-14 

SSD-CR-65-206 
00 

O 

N 

O 

4.4 

4.4 

t 0. ., 6 sec 11, - 2.2 deg 

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V-15 

SSD—CR-65-206 

0.0 

LI4 

00 

k.o 

It •>. 

<1 -.4 

o 

w 00 

ca " 

00 

ca. 4-1 

0 a RI 

N 

4.1 

t = 1.0 sec A* sa 5.7 deg 

t44 

0 

r-t 

CO 

tr) 

0 

4.1 

a) 

0 Staging Rockets - 4 Forward, 

00 

0) 

oo 

4.4 

Ce) 
00 

ro 

cu 

rti 

ers 

Co 

U 

4.1 

t = 0.96 sec A* = 2.2 deg 

t = 0 sec 1* = 0 deg 

t = 0.56 sec A* = 0.9 deg 

OD 

Lw 

00 

CO 

J.J 2 CS-156 SRM 

1/40 

00 

4,4 

44 

Staging Rockets -5 Forward, 4 Aft 

r. 

CS 

ti 

00 

a) -0 

Co

tri 

1/40 

t = 1.76 sec A* = 4.2 deg 

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V-16 SSD-CR-65-206 
b0 00 

0 al 

"Ci 

rl c‘i 

a a 

<3 <1 

U 0 

ct 

a a 

t 

U U 

it 0.64 secA4r = 0.4 deg 

sec Air = 1:0 deg 

U 

t=1.44 secAlit=1.4 deg 

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V-17 

SSD-CR-65-206 

bo 

4.) 

0 

U 

O 

a 

U 

t st 0.64 secA* at 0.6 deg 

td) 

r. 

U 

U) 

8 

sec 

 M

i 

= 

 1.

5 deg 

3 

ba 

of a1 

cx) 

0-4 4-1 

a a 

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oi 

N 

4 

cc) 

cs1 

U U 

P 

0 

4 U U0 

bo 

GL U 

co 

U a 

U) 

4 

e'S 

U 

4) 

0 

0 U

ba .,-1 
aging Rockets -5 Forward, 3 Aft 

ti) 

Fig. V-6 (cone') 

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V-18 

SSD-CR-65-206 
V-19 

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bSD-CR-65-2O6 

3 CS-156 SRNStaging 

b.0 
t 

so 0.6 sec &4,. 0.43 deg 

(d) Staging Rockets - 4 Forward, VD 

t 

V't 1.4 sec Cl4r= -0.30 deg 

t = 1.8 sec Ail- -1.47 deg 

t 

- 0 sec Aq ■ 0 deg 

t 

■ 1.0 sec At- ■ 0.49 deg 

Fig. V-7 (cont) 

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V-20 SS13-CR-65-206 

(c) Staging Rockets - 5 Forward, 

60 to 00 

w w w 

Ts no 

ssa co ran 

eu 

e•-.1 

0 

I 

1.) 

a 

I 

,4 

41 

02 

0 

• 

e..1 

11 

4.) 

N 

-3 

4 

4) 

CO 

• 

1.4 

I 

4) 

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0) 

0) 

• 

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471 

144 

02 

0. 

In 

r-4 

z 

I 

lei 

1 

..-1 

0 

g 

w 4 

4.4 

0 

rI 

41 
Fig. V-7 (concl) 

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17-21 

' SSD-CR-65-206 

4J 
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V-22 'SKY-CR-65-206 

1. 7 seg-120 

Figure V-5(a) shows the 7 seg-120 nominal motion with five 

staging rockets forward and aft, and Figure V-5(b) shows four 

staging rockets forward and aft. Both configurations give a noseout motion similar to the motion of the solids of Titan IIIC. 

The relative motion of the 4 forward/4 aft (Titan IIIC configuration) is considered adequate and is therefore recommended. The 

additional separation achieved by the 5 forward/5 aft configuration is not considered to be sufficient to warrant two additional 

staging rockets per SRM. 

Figures V-5(c) and V-5(d) show the relative motion with one 

staging rocket failure forward. Satisfactory separation is demonstrated in the presence of adverse plume loads. 

Figures V-5(e) and V-5(f) show the relative motion with one 

staging rocket failure aft and maximum residual SRM thrust on the 

left SRM. Residual thrust rotates the aft end of the SRM toward 

the core at separation and combines adversely with an aft staging 

rocket failure. Adequate separation is still achieved. 

2. 2 CS-156 

Figures V-6(a) and V-6(b) show that a 5 forward/4 aft balance 

of staging rockets is necessary to achieve the desired nose-out 

motion for this configuration. Figures V-6(d) thru V-6(f) show 

one staging rocket failure forward or aft still result in satisfactory clearance. 

3. 3 CS-I56 

A staging rocket configuration of 5 forward/4 aft (Fig, V-7(a) 

and V-7(b)] is desirable and is recommended for this configuration 

also. Again, staging rocket failures forward or aft combined 

with adverse conditions in Fig. V-7(c) thru V-7(f) demonstrate 

satisfactory clearance. 

In each of the SRM separation regimes, the importance of core 

plume force impingement on the SRM is apparent. More exact determination of the plume effect from wind tunnel tests of the selected 

MOL booster may require minor readjustment of the staging rocket 

balance. 
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VI, SOLID ROCKET MOTORS 

As a part of the performance improvement study, Martin Company 

was directed to obtain information from solid rocket motor (SRM) 

manufacturers and determine the optimum SRM design for the six vehicle configurations under consideration. The six configurations 

are defined as the core vehicle with and without transtage with 

each of three SRM options -- 7 segment, 120-in., SRMs (7 seg-120), 

2 center segment, 156-in, SRMs (2 CS-156), and 3 center segment, 

156-in. SRMs (3 CS-156). If insufficient data were received from 

the SRM manufacturers, Martin was to generate the data required 

to complete the study. 

The basic objective of the Titan III/MOL SRMs is to place the 

heaviest payload possible in an 80-n-mi polar orbit with launch 

from the Western Test Range (WTR) and within the constraints and 

ground rules specified in the Titan III/MOL Statement of Work. 

The objective of this report is to present the recommended SRM 

configurations and results of analyses leading to those recommendations, 

The following SRM manufacturers supplied SRM design data for 

use in the study: 

1) United Technology Center (UTC) supplied designs for 

the 7 seg-120 and 3 CS-156 SRMs; 

2) Lockheed Propulsion Company (LPC) supplied designs for 

the 2 CS-156 and 3 CS-156 SRMs; 

3) Thiokol Chemical Corporation (TCC) supplied a design 

for the 2 CS-156 SRM; 

4) Aerojet-General Corporation (AGC) supplied a design 

for the 2 CS-156 SRM. 

The following ground rules supplied to the SRM manufacturers 

at the start of the MOL 60-day study (2 July 1965) were intended 

to specify only general constraints to allow the SRM manufacturers 

the most latitude possible in optimizing the SRM system design: 

1) SRM attach locations, 

a) Aft - Station 1226 (all motors), 

b) Forward - Approximately Station 504 (2 CS-156), 

c) Forward - Approximately Station 250 (7 seg-120 

and 3 CS-156); 
2) Use proven state of the ar_t_for SRM design - Class 2 

propellant, existing case technology; 

3) Specific design optimization to be established by the 

SRM manufacturer based on the requirements below, 

a) (T/W)liftoff 

1.6 g for 5 sec (1.8 g is desired), 

b) qmax 900 psf nominal (+3a 5. 1000 psf), 

c) ((Thrust - Drag)/Weightlmax 3.2 g after 100 sec 

burning time, 

d) Tailoff, 

Thrust differential between two motors not to 

exceed 280,000 lbf at 1.7 g, 

Rate of thrust decay, 10%/sec of web burnout 

vacuum thrust, 

e) Thrust vector control (TVC) system, 

Type - Not specified, 

Control capability - ±8% vector deflection, 

Slew rate - 30 deg/sec maximum, 10 deg/sec 

minimum, 

The TVC requirements were adjusted on 30 July to 

reflect ±6.1% vector deflection required for the 

2 CS-156 and 3 CS-156 SRMs, 

f) Nozzle, 

Area ratio - Not specified, 

Nozzle/motor cant angle - 6 deg, 

Exit diameter - 135,8 in. maximum, 

Nozzle exit plane station - Station 1450 maximum (to ft of nozzle exit), 

SRM/pad support - Stations 1326 to 1348, 

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g) Mission success/crew safety, 

Thrust termination system is required (desired 

thrust termination curves supplied the contractors on 30 July 1965 are shown in Section C of 

this chapter) 

Redundancy is to be provided wherever practical 

in the motor design, 

Safety factors 1.25, 

Burnthrough sensors, 

3-sec minimum warning time from 0 to 

40 sec burning time and from 90 sec 

to burnout, 

6-sec minimum warning time from 40 to 

90 sec burning time, 

h) Ignition reliability - Maximum 200 ppm failures, 

i) Range safety destruct system (RSDS) is required, 

j) Inadvertent separation destruct system (ISIS) is 

required. 

Adequate information concerning the core performance, weights, 

and time sequence were also supplied to the SRM contractors. 

UTC was also working on a funded study for Air Force/Aerospace 

that supplied some additional constraints for the 7 seg-120 SRM. 

The most important of those contraints are: 

1) Meet performance constraints with minimum changes on 

the Titan IIIC 5 segment motor; 

2) Use existing Titan IIIC TVC system; 

3) Use existing Titan IIIC thrust termination ports; 

4) Use existing Titan IIIC case design. 

Some of these additional constraints were reflected in the designs 

submitted for the 2 CS-156 and 3 CS-156 SRMs. 
D`R-65 

A. REFERENCE CONFIGURATIONS 

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VI-4 

This section describes the SRM configurations used for the vehicle loads analysis and TVC system analysis, and the SRM configurations selected for SSD-CR-65-204A* and SSD-CR-65-204B.t Differences between the configurations result from the fact that the 

loads and TVC analyses were conducted on preliminary SRM data. 

The SRM criteria documents contain requirements and data resulting from the entire 60-day MOL study and reflect the preferred 

design. The general SRM configuration is shown in Fig. VI-1. 

Basic components are as shown and apply to all motor designs received. Each configuration contains the following common systems: 

1 A liquid inject TVC (LITVC) system using nitrogen 

tetroxide(N20 as the injectant fluid and gaseous 4) 

nitrogen N2 as the pressurizing gas. The N204 is 

injected into the SRM nozzle exhaust stream through 

24 injectant valves, Each configuration makes maximum 

use of existing Titan IIIC TVC system components. TVC 

injectant load and flow requirements are as specified 

for each configuration. Since the TVC sizing analyses 

were not completed in time to rerun all cases for performance evaluation, two values are listed for TVC system weights. The reference configuration values are 

those submitted by the SRM contractors, and the required values are those resulting from the TVC sizing 

studies. Additional analyses are required to show 

final adjustments in performance as a function of the 

final system weights. TVC system weights obtainable 

are less than those initially used, so payload performance will improve slightly when the final analysis 

is complete, 

2) A thrust-termination system composed of two blowout 

ports is located in the head end of the SRM. Thrusttermination stacks through the forward fairing are 

oriented in the pitch plane and at a 45-deg angle to 

the SRM centerline. Ports are opened on command by 

initiating shaped charges located around the port circumference. 

3) A separation system consisting primarily of small staging solid rocket motors located forward and aft on the 

SRM as shown in Fig, VI-1. The number of staging motors 

required for each configuration is discussed in Section 

C of this chapter. 

156-Inch Solid Rocket Motor Desi n Constra 

SSD-CR-65-204A. Martin Company, Denver, 

Colorado, August 1965. 

tTitan III/MOL 156-Inch Solid Rocket Motor Design Constraints 

(Three Center Segments). SSD-CR-65-204B. Martin Company, Denver, 

Colorado, August 1965. 
tO 

tr) 

r....1 

cri 0 

'C 

WI 

r4 

I 

0 

1146.5 1410.5 

, 19% Submerged 17% Submerged 

= 10.2:1 = 8.0:1 

Maraging Maraging 

1,948,560 2,537,100 

0.881 0.888 

' 447 x 106 588 x 106 

Parameter 7 seg-120 

Overall Length (in.) 1311.5 Nozzle Type - Conical 0% Submerged 

E = 9.6:1 

Case Material - Steel D6-AC Total Weight (Lb) 1,364,626 Overall Mass Fraction 0.864 

Total VacuumImpulse (lb-sec) 308 x 106 

^s4 

Stage 0 Design (Loads Reference Configuration) 

VAT,..15 

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VI-5 §5D-CR-65-206 
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VI-6 

seg-120 

The 7 seg-120 SRM reference confi - ation is a T 

of the Titan IIIC 5 seg-120 SRM. The reference configuration 

motor physical characteristics are shown in Fig, V1-2 and Table 

VI-1, Motor performance data are shown in Table V1-2 and Fir. 

VI-3, TVC side force characteristics, obtainable thrust vector 

deflection, slew rates, and related performance etnraet:eri.tics 

are defined in Section C of this chapter. 

The nozzle for the 7 seg-120 configuration is similar to the 

external type used on the Titan IIIC 5 segment SRM. The throat 

diameter has been increased and the nozzle lengthened to improve 

motor performance, 

Tailoff thrust-time data used for the staging analysis are defined in Section C of this chapter. The detailed tailoff and 

burning time tolerance discussion in Section B of this chapter 

defines the analysis used to obtain these tailoff characteristics. 

The 7 seg-120 grain configuration consisting of an 8-point 

star head end, 7 cone frustum center segments with partial inhibiting, and an 8-point star aft end is shown in Fig. VI-4, 

2. 2 CS-156 

The 2 CS-156 SRM reference consignration physical charactor 

ics are shown in Fig, V1-5 and Table VI-1, Motor performance 

and ballistic data are shown in Table V1-2, The motor thrusttime history is plotted on Fig. VI-3, The TVC side force characteristics, thrust vector deflection, slew rates, and related performance characteristics are defined in Section C of this chapter. 

The nozzle used for the 2 CS-156 SRN is submerged approximately 

195 of its length into the SRM chamber. The reasons for selectg the submerged nozzle are discussed in Section B of this chapter. The tailoff thrust-time curve used for the staging analysis 

is included in Section C of this chapter, This curve ha 

adjusted to reflect Martin Company analysis as defined in Fig. 

V1-3, 

The 2 CS-156 grain design consisting of a 10-point star in 

the head end, 2 conical frustum (uninhibited) center - segments, 

ulti{:1e conical frustum a -end segment shown in Fig. 

VI-6, 
F. VI-2 7 s'g-120 $IM Reference Config 

wt 

4 —was. 

Lt. 

ma, 

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Table VI-1 SRM Reference Configuration Characteristics 

7 seg-120 2 CS-156 3 GS-156 

Case (44,487) (62,260) (82,460) 

Forward Closure 4,238 

Segments 37,170 61,076 81,157 

Aft Closure 3,119 

Miscellaneous 350 1,184 1,303 

Insulation + Liner (13,330) (14,300) (16,750) 

Forward Closure 1,145 

Segments 9,625 

Aft Closure 2,560 

Propellant (571,324) 840,000 1,104,450 

Forward Closure 35,110 

Segments 518,224 

Aft Closure 17,990 

Nozzle (9,574) 10,750 11,250 

Throat Assembly 1,563 

Exit Cone 8,011 

TVC System (26,495) (29,500) (35,900) 

lnerts 10,257 15,500 17,100 

Pressurant 1,488 

Injectant (14,750) 14,000 18,800 

Usable 12,180 

Additional Fluid 1,482 

Fill Linea + Manifold 1,088 

Thrust Terrinacion (1,311) In Case In Case 

Stacks 660 

Covers 541 

Attach Hardware 60 

Mechanism 50 

Destruct System 174 50 50 

Igniter (376) (1,000) (1,000) 

Inerts 288 825 825 

Charge 88 175 175 

Hydraulic System 221 

Electrical System 722 

Instrumentation 490 

Separation System (1,289) 1,250 1,250 

Motors 696 

Circuitry 33 

Support Hardware 560 

ilose Fairing 1,222 1,500 1,500 

Aft Skirt + Hest Shield 8,756 12,700 13,400 

Forward Ring 1,352 

External Insulation 672 

Miscellaneous 128 970 1,370 

Total 682,313 974,280 1,268,520 

Expendables 

Nozzle 

Internal Insulation 5,667 7,261) 8,510 

Propellant. 511,324 840,000 1,104,450 

Igniter 88 175 175 

TVC Fluid 12,180 14,000 18,800 

External Insulation 105 

Miecellaneous 

Total 589,354 861,435 1,131,933 

Mesa Fraction 0.864 0.881 0.888 

Motor Masa Fraction 

(SRM + TVC) 0,882 0.896 0.902 

Motor Length 

Nozzle Data 

Throat Area (in,2) 1301 1/05 1630 

Exit Area in.2) 12,490 12,291 13,121 

Expansion tlatio 9,6:1 10,2:1 8.05:1 

Nozzle Submergence (%) 0 19 17 

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VI-8 
Configuration 

2 CS-156 

1200 

157 

147,4 

1,961 

670, 

2.940 

2.888 

1,104,450 

PEAR AP 

+ AL 16.1% 

40 to 100 

10-pt Steer 

Cyl Per! 

Cyl Peri' 

7 sag-120 

920 

124,31 

115.09 

1.30 

580 

1,5388 

1,4971 

1.2858 

571,324 

PEAR + AP 

+ AL 16.1% 

40 to 90 

8 Spokes 

Cyl Perf 

B Spokes 

Performance (Nozzle Centerline). 

1200 

142,0 

132.0 

1,659 

740. 

2.238 

2,187 

840,000 

PEAR + AP 

+ AL 16.1% 

' 40 to 100 

10-pt Star 

Cyl Perf 

Cyl Perf 

Maximum Expected Operating Pressure 

MEOP (psis) 

Action Time AT (sec)* 

Pressure Burning Time, P8T (sec)* 

Average Vacuum Thrust (10

6 

Average Aft Pressure (psis)* 

Impulse, AT (108 lbf-sec) 

Vacuum 

Sea Level 

Impulse, PVT (108 -sec ) 

Vacuum 

Sea Level 

Propellant 

Propellent Weight (lb) 

Composition 

Operating Temperature Range (OF) 

Configuration 

Forward Segment 

Center Segments 

Aft Segments 

Web Thickness (in.) 35.5 50.0 50,0 

Burn Rate, 100 psi, 60°F (in./sec) 0.356 0.41 

Burn Rate Exponent 

l

l 0,260 0.4 0,40 

Specific Impulse, Standard (lbf-sec/lbm # 248 248 248 

Temperature Coefficient of Pressure 

(%/°F) 0.130 0.11 0,11 

Ratio of Specific Heats 1.180 1,17 1,17 

Density (16/cu in.) 0.063 0.065 0,065 

Characteristic Exhaust Velocity (fps) 5,170 5,200 5,200 

Molecular Weight of Exhaust Gas 19,9 19.9 19,9 

Molecular Weight of Exhaust Products 26.4 27.5 27,5 

Motor Performance (without Transtage) 

Payload 28,306 35,030 42,109 

Max Q 913 927 904 

Aerodynamic Beating Indicator 96,9 x 10 6 101,2 x 106 89,3 x 10 

{Thrust - Dra. Max during First 5 sec 1,76 1,66 2,0 

Weight 

10% P to 10% P 

max max 

10% P to Web Burnout 

max 

*Averaged Over PBT 

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VI-9 

SSD-CR-65-206 

Table VI-2 SRM Performance Characteristics 
Thrust vs Time, Reference 

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V1-10 

SSD-CR-65-205 
Section B-B Section A-A 

eference Configuration 

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VI-13 
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, VI -14 
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vI-15 

3 CS-156 

The 3 CS-156 SRM reference configuration physical characteristics are shown in Fig, VI-7 and Table VI-1, Motor performance 

and ballistic data are shown in Table VI-2, and the thrust-time 

curve is as shown on Fig. VI-3, The TVC side force characteristics, slew rates, and obtainable thrust vector deflection characteristics are defined in Section C of this chapter. The nozzle 

used for the 3 CS-156 SRM is submerged approximately 17%. Section 

B of this chapter contains the discussion concerning selection of 

the submerged nozzle, The tailoff thrust-time data used on the 

staging analysis are shown in Section C of this chapter, and, as 

in the case of the 2 CS-156 tailoff, the adjustments made to the 

data are explained in Section B of this chapter. 

The 3 CS-156 grain design consisting of a 10-point head-end 

star, 3 variable conical center segments (unrestricted), and a 

variable conical aft segment is shown in Fig. VI-S. 
Reference Configuration 

H 

i•■ 

4-1 t);) 

00 0 

Pq 

O 

• 

ca 0 

91-4 CIJ 

J-k 

W E-1 CU 

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. -CR-65-206 

B. SRM CONSTRAINTS REPORT CONFIGURATION 

As a result of the 60-day MOL study, documents have been published describing the recommended 2 CS-156 (SSD-CR-65-204A) and 

3 CS-156 (SSD-CR-65-204B) . Since UTC is designing the 7 seg-12.0 

SRM under an Air Force contract, no constraints document was 

written for that configuration. However, the 7-segment configuration used in the loads study (see Section A of this chapter) is 

the 120-in, SRM configuration recommended to be used in future 

120-in. SRM MOL studies, 

Some differences exist between the configurations discussed 

in the loads reference configurations section and the motors discussed in this section. The loads configuration was based on 

preliminary 2- and 3-segment data received from LPC. 

The SRM configurations described below represent the application of the latest available data from the participating contractors together with the Martin Company's analysis of the data and 

judgment establishing motor design parameters well within the 

state of the art. 

The physical characteristics of the 2 CS-156 and 3 CS-156 SRM 

configurations recommended by Martin and described in the referenced constraints documents are shown in Fig. VI-9 and VI-10 

and Table VI-3. Motor performance data (discussed below) are 

noted in Table VI-4, and the desired thrust-time histories are 

shown in Fig, VI-11 and VI-12, 

The recommended configurations were selected after rev ew•and 

analyses of all data submitted by the participating SRM contrac—) 

tors and are judged to be well within the present SRM technology 

and state of the art. The items discussed in the following subsections were judged to be of major significance and are presented 

to substantiate the recommended SRM configurations. 

PRIK3uitu 
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VI-20 

SSD-CR-65-206 

Table VI-3 Recommended 156-in, SRM Physical Characteristics 

2 CS-156 3 CS-156 

Total Métor Length (in.) 11,146 1,396 

Maximum Nozzle Exit Diameter (in.) 135.8 135.8 

Nozzle Cant Angle (deg) 6 6 

Maximum Nozzle Submergence (p.) 25 25 

Nozzle Expansion Ratio 10:1 8:1 

Approximate Propellant Weight (lb) 840,000 1,095,000 

Usable TVC Fluid Injection Weight (lb) 9,988 11,050 

Minimum Motor Mass Fraction 0.885 0.890 

Estimated Total SRM Weight (lb) 950,000 1,232,000 

Table VI-4 Recommended 156-in. SRM Performance Characteristics 

2 CS-156 3 CS-156 

Ballistic 

146 

137 

2.22 x 108 

2,18 x 108 

1.6 

263 

Class 2 Composite 

PBAA/AN-AP 

87 Maximum 

Center Segments 

Interchangeable 

263 

156 

148 

2.90 x 108 

2,85 x 108 

1.6 

Class 2 Composite 

PBAA/AN-AP 

87 Maximum 

Center Segments 

Interchangeable 

Desired AT (sec) 

Desired PBT (sec) 

Total Delivered Vacuum Impulse (lb-sec) 

Delivered PBT Vacuum Impulse (16-sec) 

Initial Required T/W Ratio 

lb-sec Minimum Vacuum Isp ( lb / 

Propellant 

Type 

Solids Loading (%) 

rain Configuration 
,kkkk. 

kr 

0 

20 40 60 80 100 120 140 160 

Time (sec) 

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V1-25 

2.4 

2.2 

2.0 

1.8 

20 40 60 

Thrust vs Time (60°F) 

80 100 120 140 160 

Time (sec) 

Fig. V1-11 2 CS-156 SRM Constraints Document Configuration, Thrust vs Time 

Fig. V1-12 3 CS-156 SRM Constraints Document Configuration, Thrust vs Time 
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VI-26 

1, Nozzles 

A partially submerged nozzle (up to 25% submergence) is recommended for use with the 156-in. SRMs, This recommendation is based 

on the following facts: 

1) An increase in propellant weight and, therefore, payload is obtained in a length-limited system by using 

a submerged nozzle. Approximately 30,000 lb of additional propellant can be added to the selected configurations as a result of 25% nozzle submergence. 

2) The submerged nozzle state of the art is firmly established for the Minuteman program and is supported by 

numerous subscale firings with throat diameters up 

to 15 in, In addition, three 156-in.-diameter motor 

submerged nozzle tests are scheduled in the near 

future. LPC is scheduled to test fire two motors before January 1966, and TCC will fire one motor early 

in 1966, 

3) Martin Company tradeoff studies show that the heat 

transfer, structural integrity, material erosion, and 

char characteristics of the partially submerged nozzle 

are most acceptable for use on large SRMs. The various 

nozzle materials and properties submitted by the participating SRM.contractors are listed in Section C 

of this chapter. Martin Company recommendations are 

noted for the nozzle throat material (graphite phenolic) 

and exit cone material around the TVC injectant port 

location (silica-phenolic). Graphite-phenolic provides 

an acceptable erosion rate at the throat and, from test 

data, appears to maintain structural integrity during 

long-duration firings better than other materials. 

Most N2

0

4 

TVC experience has been with silica-phenolic 

insulation surrounding the TVC injectant ports, Its 

erosion rate in the presence of oxidizing N2

0

4 

is known 

and acceptable for long-duration firings. 

4) Because of the proven state of the art of submerged 

nozzles for smaller solid-propellant motors and because of forthcoming large motor submerged nozzle 

tests, the technical risk associated with the 156-in. 

SRM submerged nozzle is low. It is far less than the 

risk taken with nozzle development on the 120-in. 5-

segment motor at a similar program point. 
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2. TVC System 

An N2

0

4 

LITVC system is described in the SRM constraints documents and is presently the recommended system, However, if any 

other type TVC system with equal or better reliability and operating characteristics is proposed in the future, consideration will 

be given to that system. As mentioned previously, all SRM participants proposed LITVC using as much of the existing Titan IIIC 

120-in. 5-segment SRM TVC system as practical. Martin Company 

LITVC loading, thrust deflection, and slew rate requirements are 

listed in Section C of this chapter. The values listed as well 

as the recommendation for use of an LITVC system are based on 

the following: 

1) Analysis of the MOL SRM system TVC requirements show 

that the LITVC system presently used on Titan IIIC 

can meet the requirements for thrust vector deflection at maximum motor thrust, slew rate, thrust vector 

deflection during tailoff, and injectant fluid load 

requirements with minor modifications; 

2) Fourteen static tests and two flight tests of 5-segment 

120-in, motors have shown that the LITVC system is reliable and capable of being used successfully on large 

SRM systems. LITVC use on the Polaris and Minuteman 

programs has also shown considerable success; 

3) The success of LITVC on the Titan IIIC program and 

availability of the subsystem test data and common 

components indicate a very low technical risk associated with the TVC system, The minor component redesign and requalification required will not be the 

pacing item on the 156-in. SRM program. 

3. Cases 

The case material recommended for use on the 156-in. SRMs is 

250 grade 181 nickel maraging steel. The major reasons for selection of this material are: 

1) Four 156-in. SRMs using nickel maraging steel cases 

have been successfully static fired. Maraging steel 

cases have also been successfully used on the 260-in.-

diameter SRM programs; 
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VI-28 R-65-206 

2) Existing steel production technology such as vacuum 

arc remelting with controlled rates of solidification 

and homogenization of the ingot before hot working, 

and welding techniques (high rate metal deposition 

with TIG method) allow production of 250 grade maraging steel with fracture toughness of better than 100 

ksi in.; 

3) The 18% nickel maraging steel has excellent capability 

for repair and rework, resulting in a lower rejection 

rate. This is a cost advantage as well as a desirable 

aspect from a scheduling standpoint; 

4) The technical risk associated with the maraging steel 

cases is deemed to be relatively low from a performance 

standpoint. To maintain this low risk, Section C of 

this chapter describes the Martin-recommended case design criteria; however, long lead times are involved, 

and it is recommended that strong consideration be 

given to ordering material before a May 1966 program 

go-ahead. 

4. Tailoff Parameters 

Initial criteria for SRM thrust tailoff design were based on 

limiting the maximum thrust differential between SRMs during tailoff to approximately 280,000 lbf at core-engine ignition. With 

this limitation, it was anticipated that the existing Titan III 

core aft longerons would not require redesign. SRM tailoff analysis shows that it is costly from a payload standpoint, and somewhat risky from a SRM performance standpoint, to limit tailoff 

thrust differential to the low level required by the existing 

core aft longeron design. It is strongly recommended that the 

core aft longeron be redesigned to withstand approximately 500,000 

lb of thrust differential for the 2 CS-156 SRM and approximately 

600,000 lb of thrust differential for the 3 CS-l56 SRM. The following items establish the basis for this recommendation and 

specify the Martin Company design limits where applicable for the 

SRMs and as stated in Section C of this chapter: 

1) Data submitted by SRM manufacturers indicated that with 

burning time control held to within ±1% (3 sigma), the 

280,000 lb thrust differential between SRMs was not exceeded. However, after a careful review of Titan IIIC 

120-in. 5-segment SRM data, Minuteman, Polaris, and 

numerous small motor programs, burning time control 
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VI-29 

of ±2% is recognized as state of the art. Subsequently, 

the ±2% motor-to-motor burning time variation was used 

to determine the maximum force differential between 

SRMs used for the MOL study, A 3°F temperature difference between SRMs was also used in the thrust differential analysis, Section C of this chapter discusses the results of thrust differential as a function of tailoff time with both ±1% and ±2% motor-tomotor burning time variation. This plot was obtained 

by "stretching" out the SRM tailoffs defined in the 

design contraints reports until a low level of thrust 

differential was reached. It is seen that to meet 

initial 2- and 3-segment requirements, the tailoff 

time must exceed 25 sec. 

2) The 3 CS-156 SRM payload performance was computed using 

a 22.5-sec tailoff. Payload weight was decreased by 

approximately 1700 lb (or 4%) over performance using 

the recommended tailoff time of 10 sec, 

3) The technical risk associated with the burning rate 

control tolerance of ±2% and the allowable 500,000 to 

600,000 lb thrust differential between SRMs is exceedingly low. These tolerances are firmly established as 

state of the art and are strongly recommended to be 

included on the 156-in. SRM designs. 

5, Propellant 

Propellant formulation and grain design for the recommended 

SRMs will specifically be somewhat dependent on the individual 

design. However, Martin recommendations of Class 2 composite 

PBAA/AN-AP propellant and simple interchangeable center segment 

grain design are still imposed. Note that all propellants proposed by the SRM manufacturers were of the recommended type and 

are all proven state-of-the-art formulations with well-known characteristics. 

• The technical risk associated with the grain designs is very 

low. More than twenty 156-in. and 120-in, SRMs using features of 

the required grain designs have been tested with very successful 

results. 
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6, Thrust Termination 

SRM thrust-termination capability is required for the MOL vehicle, The Titan TUC method of thrust termination is presently 

being recommended; that is, by releasing two ports located in the 

motor head end. Martin analysis is continuing and will subsequently 

fix the port size range for each configuration. To maintain core 

vehicle structural integrity following thrust termination, a positive forward force of 100,000 lb must be maintained by the SRMs. 

Section C.8 of this chapter defines the SRM force required to maintain core integrity. 

C. SRM CONTRACTOR DATA 

In addition to the recommended SRM configurations described 

in Section B, several other motor designs were received and analyzed. The data presented in this section are the final designs 

received from each participating SRM contractor. Brief notations 

and explanations of Martin Company analysis are also indicated. 

1. Internal Ballistics and Propellants 

Table VI-5 contains the internal ballistics, propellant, and 

grain design data for each of the motors submitted (final design). 

Curves of thrust, chamber pressure, and weight flow are presented 

in Fig. VI-13 thru VI-17, Ignition overshoot is ignored on these 

curves, but was covered in consideration of the maximum expected 

operating pressure (MEOF). 

The propellant formulations proposed are shown in Table VI-5 

and are all composite propellants composed of PBAA or PBAN with 

aluminum and ammonium perchlorate. All propellant formulations 

are of proven quality and are well within the propellant state 

of the art, Physical, mechanical, and chemical properties of 

each propellant proposed are acceptable. 

The propellant grain designs (all similar to Fig. VI-6) are 

also simple, proven, state-of-the-art designs with acceptable 

stress levels and design shapes. Some minor variations were submitted with respect to inhibiting (see Fig. VI-4) and port shaping, but all are acceptable (provided that the related thrusttime curve is acceptable). It is a requirement that the center 

segments be identical to maintain interchangeability within each 

of the three configurations. 
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V I - 3 1 

0 

0 

Q) 

0 

E 

3-4 

0 

3-4

0 

z 

(.11 

C.) 

CNI 

5 

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L 

(11 00 

(uTed zot) aanssaza iacituptio 

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01) lsniqy tunnovA 

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a 

0.1 0 

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6 0 

0 

P. 5 

g 4 

CJ

3 

2 

0 

R-65-206 12 

11 

10 

9 

8 to 

7 e•-) 

6 

5 

4 

3 

10 20 30 40 50 60 70 80 90 100 110 120 130 

Time (sec) 

0 

0 

1 

VacuumThrust 

3 12 12 

11 

10 10 

Pressure 

0 

0- 8 

O - 7 

a 

a 

• 

6 

a 

.a 5 

1

• 

4 

2 

a 

1 

Mass Flow 

Thrust 

10 20 30 40 50 60 70 80 

Time (sec) 

0 

90 100 110 120 130 140 0

o 

Fig. VI-15 AOC 2 CS-156 Motor Performance (60°F) 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

Fig. VI-14 TCC 2 CS-156 Motor Performance (60°F) 
V1-33 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

12 

10 

60 80 

Time (sec) 

--Thrust 

12 

10 

Pressure 

a4 

• 

20 40 60 0 

Time (sec) 

100 120 140 

Mass Flow 

Fig. V1-16 LPC 3 C5-156 Motor Performance (60°F) 

Fig. VI-17 UTC 3 CS-156 Motor Performance (60°F) 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VI-34 

SSD-CR-65-206 

Table VI-5 SIAM Internal Ballistics 

Contractor Molar a:et e 

Segments 

UTC 

120 

7 

521 

156 

3 

Lockheed 

1.56 

c 

1 ckhced 

156 

3 

Thiokol 

156 

:ter; el: 

136 

Nozzle Pate 

Throat Area, Initial (eq in.) 1,301 1,768 1,205 1,630 1,286 1,419 

1,736 

throat Agee, Final (sq in.) 1,374 1,882 1,302. 1,740 1,538 

Exit Area, Initial (sq in.) 12,490 14,144 14,000 14,000 7,724 11,310 

Exit Area, Final. (sq in.) 12,650 14,450 14,080 14,000 7,724 11,310 

17.5 Expeneion Cone Half-Angle (deg) 15.0 15.0 17.5 17.5 17,5 

7,0 

Expanainn Ratio, Initial 9.6 8,0 11.6 8.58 6.0 

1.31 Poet-to-Throat Area Ratio 1.32 1.49 1.9 1,4 1,78 

Ptrformnnec (Ntiltle) 

NEOP (pal) 920 1050 1189 1189 1400 927 

127 

Action Time (AT)(sec)* 124,31 151.39 150 157 142 

Pressure Burning Time (PBT)(sac)t 115.09 141.39 137.25 147.4 129 118 

Initial Vacuum lltrust (10' 110* 1.61 1.93 1.85 2.725 1.48 1.50 

„ \** 

Average Vacuum Thrust (10 ID6 

f initial «ft PI-eat/ore (psia)e 

1.30 

550 

1,82 

669 

1,618 

825 

1.954 

825 

1.10 

921 

1.858 

60° 

Average Aft PrO5Vre (Pah)** 

impulae, AT (108 Ibi -aec) 

580 605 818 696 770 667 

Vacuum 1,5388 2.6025 2.272 2,932 2.234 2-25 

Sea Level 

2.0579 

Impulse, PET (108 141-aec) 

Vacuum 1.4971 2.5710 2.221 2.880 2-176 2.1896 

Sea Level 1.2858 2.2768 2,0080 

0.536 

Mane Fraction 0.864 0.878 0.884 0,893 0,867 

0,88 

Banjo Motor Efficiency 0.96 0.96 0.96 0.96 0.96 

Propellant 

Propellant Designation UTP 3001 UTP 3001 LPC 580A LPC 580A TP H1011 A148-3105 

Propellant Weight 511,324 979,187 842,113 1,112,000 884,553 685,900 

PBAN + AP 

Composition PBAN + AP 

+ Al 16.1% 

PBAN + AP 

`A1 16,1% 

PBAN + Al' 

+ Al 16.17, 

MAN + AP 

+ Al 16.17. 

1144/AN 

+ ,r + .11 16,17, t• Pe 0 -' ,0 157. 

Operating Temperature Range ( ° F) 40 to 911 to 90 40 to 100 40 to 100 40 to 90 

Configuration 

Forward Segment 8 Spokes 10-pt Star 10-pt Star 10-pt Star 4 Spokes Conocyl 

Center Segments Cyl Pea Cyl Pea Cvi Perf Cyl Perf Cyi Per( Cyl Perf 

Peri 

Aft. Segment 8 Spokes Cyl Perf Cy1. Perf Cyl Perf Cyl. Serf Cyl 

51 

Web lliickneas (in.) 35.5 49.0 50,0 50.0 48 

0.474 

Born Rate, 1000 psi, 40°F (111./sec) 0.3.56 U.385 0.41 0.331 0.392 

0.33 

Burn Rate Exponent , 

0.260 0.240 0,4 0,40 0.21 

Specific Impulse, Standard (16E -see 

/56 1 m7 

248 248 248 248 246 244.6 

Temperature Coefficient of Preen. 0.16 

(74/7F) 0.130 0.130 0.11 (1.11 0.10 

Temperature Coefficient of Burn. 0.107 

Sate (37°F) 0.096 0.099 0.07 0.07 0,08 

1-18 

Ratio of Specific Bests 1.180 1.180 1,17 1.17 1.18 

0.0435 

Density (16/60 in.) 0.063 0.0635 0.065 0.065 0.064 

Characteristic Exhaust Velocity (fps) 5,170 5,160 5,200 5,200 5,180 5,120 

5,633 

Adiabatic Flame Temperature (*F) 5,700 5,674 2,87U 5,870 5,790 

Molecular Weight of Exhaust On 19.9 19.9 19.9 19.9 18.7 

, 28.8 

Molecular Weight of Exhaust Products 26.4 26.4 27.5 27.5 26.4 

Exhaust Cas Composition 

Exit Plant 

112 30.3 30.3 30.0 30,0 27.9 25.2 

H20 11.3 11.3 11.3 11.3 14.1 17.2 

Ii 0.2 0.2 0.6 0.6 0.2 0.4 

25.0 

CO 25.4 25.4 22.0 22.0 23.0 

2.6 

002 1.9 1.9 1.9 1.9 2.3 

Na 7.8 7.8 8.8 8.8 8.4 8.7 

/111 15.0 15.0 15.7 15.7 15.9 16,4 

8.0 

A1207 7.9 7.9 9,1 9.1 8.0 

CL 0.1 0.1 0.3 4.3 0.) 

...... 

0.2 

.. 

Fe -- 

__ 0,3 0.3 

*101; Pm.. to 10% Pm.... 

9107, P,,..„ to web burnout. 

0/let/Bored at time when P .. 757. Pm,,,,, 

**Averaged over VAT. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VI-35 

2, Performance Tolerances and Transients 

Basic performance parameters and their associated tolerances 

are shown in Table VI-6. The values recognized as state of the 

art and recommended for use are those values listed in the "Martin 

Recommended" column. 

Tailoff variations submitted by each SRM contractor are shown 

in Table VI-7 in conjunction with Fig. VI-18. Also shown in Table 

VI-7 are the tailoff variations resulting when the recommended 

state-of-the-art tolerance values are used. 

Figures VI-19 thru VI-21 present the reference configuration 

tailoff curves used for loads and staging calculations. Figure 

VI-19 (7 seg-120) is as presented by UTC. Figures VI-20 and 

VI-21 (2- and 3-segment reference configurations) use the tailoff shape presented by LPC, but incorporate Martin tolerances 

for burning times. Figure VI-22 presents the TCC 2-segment tailoff, which incorporates a long tailoff time with ±2% time tolerances in an attempt to limit thrust differential to a maximum of 

300,000 lbf. Note that the longer burning time TCC motor produces a lower payload by approximately 2000 lb than the comparable reference configuration 2-segment motor. 

In an attempt to define payload loss with increased tailoff 

burning time, the reference 3-segment configuration thrust time 

tailoff was increased from 10 to 22,5 sec duration (Fig. V1-23), 

Total vacuum impulse was held constant. The payload lost was 

approximately 1700 lb, or 4%. 

3, Ignition 

Ignition variations submitted by each SRM contractor are shown 

in Table VI-8 in conjunction with Fig. VI-18. The igniter designs 

and associated ignition transients submitted by the SRM contractors 

were not all acceptable. 

Figure VI-24 indicates the relative position of each igniter 

proposed for the parameters defined. Additional constraints of 

bore pressure and heating rate were also considered, However, 

an igniter producing the transient recommended (Table VI-8) was 

bracketed by submitted designs and is well wlthin present industry igniter design capability. To minimize risk in ignition and 

igniter design, it is strongly recommended that an SRM head-end 

mounted pyrogen ignition system be used. A summary plot of ignition curves received is in Fig. VI-25. 
SSD-CR-65-206 

Table V1-6 SKR Performance Tolerances (60°F) 

Martin 

Recommendation 

7 seg-120 2 CS-156 3 CS-I56 

urc LPC ICC ACC UTO LPC 

Vacuum Total mpulse 

PBT (7,) 1.25 N/A N/A N/A N/A N/A N/A 

Vacuum Total Impulse, 

AT (%) 1.0 1.0 0.69 1.0 0.81 1.0 0.69 

Specific Impulse (%) 0.7 0.7 0.45 0.70 0.13 0,7 0.45 

Action Time (AT)(7,) 3,0 3,43 N/A 2,0 N/A 3,09 N/A 

Pressure Burning Time 

(PBT)(%) 2,0 2,16 0,96 2.0 2,0 1,95 0,96 

Ignition Interval (%) 15.0 9.1 6.0 11,1 20,8 4,9 6.0 

Variation in Thrust 

before Tailoff (7,) 3,0 4.04 1.3 2.2 2.38 4.04 1 , 3 

All values are t3 sigma. 

Table VI-7 Tailoff Parameters 

Parameters Submitted by SRM 

Contractors 

7 seg-120 2 CS-156 3 CS-156 

UTC LPC TCC AGC UTC LPC 

Atb 

Max (sec) 4.98 2.50 5.12 4.72 5,50 2,82 

At

a 

max (sec) 8,52 2.50 5,12 9,34 2.82 

AF max (103 lb

f 

) 340 280 300 500 280 

Max Slope (%/sec) 8.1 8.0 8,8 9.0 11,5 10.0 

Parameters Computed by Martin 

with ±2% Burning 

Time Tolerance 

Ltb 

max (sec) 3,40 4.30 4.20 3,40 4.00 

At

a 

max (sec) 6,40 6,00 6,00 6.20 4.60 

AF max ( io lbf) 340 500 340 520 600 

Initial Criteria: 

AF max (All Systems) = 300,000 lbf max and 280,000 lbf at 1,7 g. 

Slope = 10%/sec of Web Burnout Thrust 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VI-36 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

Ignition and Tailoff Parameter Definitions 

VI-37 

(%) wnwlxvN 
0 0 a o 0 

) lenagy tunnaA 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

V1-38 

Fig, VI-19 Tailoff Vacuum Thrust Curves, 

0 

CN 

n-4 

1 

of 

f•-•• 

SSD-CR-65-206 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

V1-39 
Thrust Diffe

rential 

S 

z 

400 

300 

200 

0 10 20 30 

Tailoff Time (sec) 

SSD-CR-65-206 

3-Segment SRM Payload 40- —4061A Loss Associated with 

12.5 sec Increased 

Tailoff is Approximately 1700 lb 

_ 

\\ 

\\ 

\ \\ 

\\ 

Limit 

'A\ 

Ns. \ 

---7-Segment 

at 1.6 g \ N. 

-c-- 

N 

...„,, CS-156 

3 CS-I56 

----I--- 

seg -120 

2- aiid 

___ —....., 

3 CS-156 

--- ...-- 

Limit at 1.85 g ------- 

--7 

Legend: 2 CS-156 

±1% Variation on Motorto-Motor PBT 

±2% Variation on Motorto-Motor PBT 3 CS-156 

Fig.,VI-23 Maximum Thrust Differential vs Tailoff Time 

700 

600 

100 

0 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VI-40 

4.4 500 

r-1 
"",4' 200 -- 

5 

150 

rn 

'D 

• 

100 

4.1 

0 

0 

0,08 0.16 0 24 2 

Marginal 

or Missfire 

Good Ignition 

300 

250 

TCC 

156 

AGC 

156 — 

IPCry 

156-3 

Previous Thiokol 

Motors Contained 

by Band 

0.10 0.20 0.30 0,40 0.50 0 60 0.70 0.80 

Burn Time (sec) 

Maximum Vacuum Thru st (7) 

70 

60 

50 

40 

30 

20 

10 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VI-41 

Fig. VI-24 Relationship of Igniter Discharge to Ignition Delay 

Fig. VI-25 Nominal Ignition Transient, Large SRMs 
2 CS-156 3 CS-156 

LPC TCC AGC C LPC 

6.4 

0,30 

8,0 

0,50 

80 

275 ± 31.5 

430 

20 

3c, 

MS 

ti 

MS 

t (100%Y), 

MS 

80 ± 15 

275 ± 25 

430 

150 95 

180 ± 10,8 565 ± 63 360 ± 15.2 310 ± 15,2 

600 

68 

180 ± 10.8 

450 

68 140 

450 1000 700 

Force, 7. 

Maximum at 

Time (sec) 

17.5 

0.250 

9.2 

0.220 

27,8 30.2 

0,475 0.270 

Martin Recommendations 

td, MS 

t. ± 30, 

1 

MS 

t (100%Y), 

MS 

L Force, % 

Maximum 

75 

215 ± 25 

400 

25 

See Fig. VI-18 for symbol definitions. 

Contractor7 seg-120 Submitted 

Parameters UTC 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VI-42 

Tab e VI- ition Pa mete 

TVC System 

All contractors proposed an N204 LITVC system to meet the defleCtion angle requirements of 8% (side force)/(axial force) ratio 

for the 7-segment motor and 6.110 (side force)/(axial force) ratio 

for the 2- and 3-segment motors. Each proposed system made maximum 

use of the existing UTC TVC system components. Table VI-9 gives a 

brief summary of the contractor-supplied data and Martin recommendations for detailed portions of the system. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VI-43 

Based on static test and flight experience, the N204 LITVC 

system is recommended for use on the MOL/Titan III system. It 

is recommended that the N2

0

4 

prevalve be retained in the TVC system until additional analyses have been conducted. 

The TVC system requirements (weights and angles) noted in 

Table VI-9 are the result of six-degree-of-freedom trajectory 

analysis using the reference configuration motors with TVC side 

force characteristics as shown in Fig. VI-26 thru VI-31. These 

side force data were based on previous test results and were adjusted to match motor design conditions. Figure VI-32 shows the 

number of valves (103 and 143 lb/sec size) required to obtain the 

thrust deflection angle noted, It can be seen that 24 valves are 

required for each of the three reference configurations. 

5. Nozzles 

Two basic nozzle designs were proposed by the participating 

SRM contractors. UTC proposed an external nozzle configuration 

for both the 7 seg-120 and the 3 CS-I56. The 7 seg-120 nozzle 

is basically the same as the nozzle used on the Titan IIIC 5 seg120 motor. The 3 CS-156 UTC nozzle is a scale-up of the 5 seg120 Titan IIIC SRM nozzle. All other contractors submitted partially submerged nozzle designs ranging from 17% submergence 

(LPC) to 45% submergence (TCC). The percentage of submergence 

is a function of both the nozzle area ratio and TVC injectant 

point. For a nozzle area ratio of 8:1 and injectant point area 

ratio of 3,5:1, the maximum submergence is approximately 25%. 

Table VI-10 shows pertinent nozzle parameters as proposed by 

the SRM contractors. Where applicable, Martin recommendations 

are also noted. 

6. Motor Cases and Attachments 

Table VI-11 lists the proposed design criteria for the 7 seg120,,2 CS-l56, and 3 CS-156 SRM motor cases and attachment structures (to the vehicle core and the TVC fluid tank). 

Because of stringent program schedule requirements, these criteria represent conventional design methods using existing tooling 

and technology. The newer methods offering potential improvements 

will present long lead times and higher overall costs. However, 

as the program develops, it may be possible to incorporate them 

into the later motor cases following intensive evaluation. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VI-44 

Table VI-9 TVC System Requirements 

Martin Company Requirements 7 seg-120 2 CS-156 3 CS-156 

TVC Injectant Fluid Required (Usable) (lb) 6,248 9,988 11,050 

Number of Valves 24 24 24 

Flow Rate of Valves (lb/sec) 103 143 143 

Minimum Slew Rate (deg/sec) 10 10 10 

Maximum Required Thrust Deflection at 

Maximum Motor Thrust (deg) 3.0 3.0 3.0 

Maximum Required Thrust Deflection 

during Tailoff (deg) 6,0 6.0 6.0 

SRM Contracter- 7 seg-120 2 CS-156 3 CS-156 

Supplied 

Parameters UTC LPC TCC AGC UTC LPC 

Nozzle Area Ratio 

at Injection Point 2.94 4.0 2.7/3.0 3.5 3.5 3.0 

Injector W (lb/ 

sec) 103 142 155 143 142 

Number of Valves 24 24 48 24 24 32 

Tank Operation P 

c 

(psi) 

750 750 750 750 800 750 

Usable Fluid (lb) 12,180 14,000 11,000 25,627 13,650 18,800 

Maximum Angle Required (deg) 4,07 5,0 7.1 6.1 4.0 4.9 

Slew Rate, Minimum 5 at 10 at 6.8 at 

(deg/sec) 6 deg 6 deg 4.58 

deg 

Slew Rate, Maximum 35 at 30 at 20.3 at 

(deg/sec) 0 deg 1.5 

deg 

2.0 

deg 
Motor 

Vacuum 

Thrust 

(106 lb) Fo (Psia) 

1.56 (700) 

1.34 (600) 

1.03 (450) 

0.48 (200) 

V1-45 

SSD-CR-65-206 

2 4 6 8 

Command Voltage (v) 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

500 

Quadrant Side FlowRate (1b/sec) 

400 

300 

200 

Side Force vs Voltage 

0 Voltage in 

Adjacent 

Quadrant 

Fig. VI-26 TVC in UTC 7 seg-120, 

600 

100 

4 6 8 10 

Command Voltage (v) 

Fig. VI-27 TVC in ,UTC 7 sagp.1414% Flow Rate vs Voltage 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VI-46 

nd Voltage (v ) 

2 C5-156, F1aw Race vs Voicage Fig. V1-29 TVC in 

0 8 0 0 0 

SSD-CR-65-206 

01 09 ao4oa *pis, 

TVC in LPC 2 CS-156, Side Force vs Voltage 

to 

to 

(.e/qt) ei ,1011 opTS 
VI-47 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

TVC inLPC 3 CS-156, Flow Rate vs Voltage 

0 0 0 0 

(0"/41) al1711 moTa olns luvaplinb 

8 8 8 

r-1 

TVC in LPC 3 CS-156, Side Force vs Vo 

ea 000 

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0 (7 0 

h 41 .0 CO 01 0 co 

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(ca) ao.zoA apis ltrinpvnb 

SSD-CR-65-206 

N 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VI-48 SSD-CR-65-206 7 MEM I_ - i I 

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NRO APPROVED FOR 

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CO .0 44 ti 

0 0) 

4.4 ". 

<49 z' 

4, 

0 N 

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00 

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0 0 

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400 a4 

Tab leVI- 10 Nozz le Characteristics 
1, Case and attachments for 7 seg-120 SRM to follow Titan IIIC 

most recent practice, 

2, Case 2 CS-156 and 3 C5-156 SRMs 

a Material 18% Ni maraging steel 

Grade 250 

Fracture toughness 100 ksi 

b. Fabrication 

Material yield strength, 

0.2% offset, ksi 230 min, 

Factor of safety to 

minimum yield 1,25 

Design pressure MEOP x 1,25 

Case weld efficiency 95% 

d, Joint type Torque and clevis with 

pin and 0-ring 

Approximate stations for attachment to vehicle core 

Forward 7 seg-120 250 

3 CS-156: 250 

2 CS-156: 504 

Aft 1226,56 

Attachment structures 

Forward structure and aft skirt Following Titan IIIC 

design practice 

Forward and aft 

closures 

Dome: Spin forgings 

Cylinder: Rolled and Welded 

Y-joint: Rolled ring forging 

Welding 

Torque and clevis 

joints 

Center segments 

Port bosses 

(Igniter, TT, 

and Nozzle) 

Rolled and welded 

Rolled ring forgings 

Rolled ring forgings 

TIC method, high-rate metal 

deposition 

c. Design 

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v1-50 

Table VI-11 Case and Attachments Design Criteria 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VI-51 

7, Insulation and Liner 

Insulation materials proposed by the SRM contractors are shown 

in Table VI-12, Figure VI-33 shows typical proposed insulation 

joints. 

Since the insulation materials proposed are very similar, 

little selection is available. However, based on previous tests, 

the preformed insulation is recommended over the mastic, Insulation joints are also similar, and either concept would be acceptable. 

Since insulation thicknesses vary with the grain design and 

burning time, they must be determined for each specific design. 

However, an insulation thickness safety factor of at'least 1.5 

is recommended. 

8, Thrust Termination 

To enhance crew safety, the capability to terminate (or reduce) SRM thrust must be available. 

To maintain core structural integrity, a net forward force of 

100,000 lb must be applied to the core vehicle after thrust termination (Fig, VI-34). Figures VI-35 thru VI-37 show the results on 

Gemini B abort when the structural requirements are met. 

Port sizing for each SRM configuration will be set (at this 

time) as a function of the structural requirements, The port sizing analysis is presently being conducted. 
SSD-CR-65-206 

Table VI-12 Proposed Insulat ionSystems 

. 

H 

E. 

to 

o 

to 

r 

a 

cr 

4.1 

to 

o •,..4 

'tot 

• er 

U 

0J 

in 

l'-• r-1 

1...4 N 

N N. 

44 . 1 1 

4 MDS 407 17 4 MDS-40723 

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H 

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to 

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NI 

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N.I 

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r rr. 

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in 

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N 

0 

N 

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N 

H f... . .1.. : 

HO 

••+-■ crl 

0 

0 

Carbon Re inforced 

Phenolicwit h 

Buns "N" Modifier 

4t1 

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EA 0 • a 

0 HO V: pi 

CV -4 

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Mast ic 

Bor ic Ac id 

Asbe stosFi lled 

AsbestosF i lled NBR and 

Mast ic 

,, 

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I ...3" 

a ,:t rir I 

(OH 

U H 

0 

m 

Z 

Center 

Segment 

0 

,-. 

..: 

1. 

•,..i w H 

r... 010 

0 0 ztr; •,.., cr 

ti 

-0 r, 

•,-r to >r, ■-.1 .4.4L r■-,i Q 

(0 (0< 

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- 

■ :4 1 .0 . 0.04 

-- r.-..: to 

■--, .--I 0 

v: 4, 01 Mast ic Boric Ac id - 

Asbe stosFilled 

AsbestosFilled 

NBR.and Mast ic 

U 

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to 

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4, 0 

.,a 

— .1) 

i: ' 0 

'0 0 

''' C.) 

0 r---, 

Asbestos Fiber 

SilicaFilled 

Butad iene 

Acry lonitrile Asbestos 

SilicaFilled 

Bun s "N" 

Mastic. 

Bor ic Ac id - 

• 

Asbe stos Filled 

Asbestos 

' Filled NBR 

and Mast ic Gen-Gard 

V 44 

Ma st ic 

Contractor 

Un ited 

Technology 

Un ited 

Tec hno logy 

Loc kheed 

Propulsion 

--, 

- 10 

0 

.i'L ,-+ 00) H ri, 

.0.0 HO 

Aeroje tGener a l 

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Propulsion 

SRM 

Con figura tion 

0 N 

.... 

(0 0) 

to 

r, 

‘0 ir: 

rr-, 

o 

0 

cn 

,0 

0", ,r....i 

til 

,..) 

N 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VI-52 
Propellant 

Boot (UF2121) - 

SSD-CR-65-206 

Z- Asbestos Filled NBR 

Fill with Zinc 

Chromate Putty \\Falloon

iller 

B

(Epoxy MicroBalloon 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 VI-53 

Propellant - 

Fill with RTV-88 

NBR-Boot 

(r) Lockeed SRM Segment Joint 

Boric Acid 

Asbestos Filled Mastic 

(a) Thiokol TU-502 Segment Joint 

Fig. VI-33 Typical Proposed Insulation Joints, 156-in. Diameter 
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RELEASE 1 JULY 2015 

V1-54 

------r-----r- ) t- _ 7 -1— 

Note: 1. 6-deg nozzle cant. 

2. Reference trajectory 

data used, with transtage. 

.o 

140 

(.1 120 

o 100 

80 

60 

40 

20 

SSD-CR-65-206 

4 

3 

2 

1 

H 

-1 

-2 

-4 

10 20 30 40 50 60 70 80 90 100 110 120 130 140 150 

Time from Liftoff (sec) 

Fig. V1-34 SRM Net Forward Thrust Along Nozzle Centerline to Maintain Total 

Thrust Along Vehicle Centerline of 100,000 lb 

11.1.11M Booster 

100,000 lb Net Forward 

Thrust after TT 

Present 1.20 -7 

1110 

TT Resign 

Oemi. ni B

. 

with Six Retro 

► Rockets 

1. UTC 7 seg-120 (reference conl 

figuration) with transtage. I 

2. Thrust termination (TT) data I 

supplied by UTC. 

1 ...L...— 

10 20 30 40 50 ,60 70 80 90 100 110 120 130 140 

Time (sec) 

Fig. V1-35 Thrust Termination Evaluation, UTC 7 seg-120 
41 

100,000 lb Net 

after 

Forward-- 

' 

Booster 

Gemini 

Si Retro 

Thrust 

B with 

1ockets 

Note: 1. LPC 3 CS-156 (reference 

configuration) with transtage. 

2, Thrust termination (TT) 

data supplied by LPC. 

10 20 30 40 50 60 70 80 90 100 110 120 130 140 150 160 

-4 

0 

4 

3 

2 

1 

-1 

-2 

3 

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RELEASE 1 JULY 2015 

VI-55 

4J50

CU 

H 

W 

E-4 

100,000 lb Net Forward Thrust 

after TT 

Gemini B with 

Six Retro Rockets 

I i 

Note: 1. LPC 2 CS-120 (reference configuration) with transtage. 

2. Thurst termination (TT) 

data supplied by LPC. 

-4 A A 

0 10 20 30 40 50 60 70 80 90 100 110 120 130 140 

Time (sec) 

Fig. VI-36 Thrust Termination Evaluation, LPC 2 CS-156 

Time (sec) 

Fig. VI-37 Thrust Termination Evaluation, LPC 3 CS-156 
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VI-56 

9. Staging Rockets 

Staging analyses show that the Titan IIIC staging rocket will 

provide the required thrust levels in all cases as defined in 

Table VI-13. The staging analysis is thoroughly discussed in Chapter V of this report. 

Table VI-13 Staging Motors Required 

7 seg-120 2 CS-156 3 CS-156 

Aft 

4 

Forward Aft 

MO*1.41.0,1 

Motors Required 

per SRM 4 

Forward 

5 

Forward 

5 

Aft 

4 4 
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VII-1 

VII. LIQUID PROPULSION SYSTEMS 

The liquid propulsion system portion of this study consists 

of comparisons between the Titan IIIC core propulsion systems 

and a core using modified liquid engines. The primary difference 

is that the use of a modified YLR87-AJ-9 engine with a 15:1 

expansion ratio is considered. 

The original work statement specified the use of a 14:1 

expansion ratio thrust chamber on the YLR87-AJ-9 engine which 

presently has an 8:1 expansion ratio. The expansion ratio considered in this study was subsequently changed to 15:1. All 

performance and weight data used in this study for the modified 

engine were therefore based on the 15:1 expansion ratio. Additional modifications including the YLR87-AJ-9 engine turbopump and 

injector have been proposed by Aerojet-General Corporation. 

The Stage II engine, YLR91-AJ-9 and transtage engines, AJ10- 

138, are used in this study with no major modifications that affect the core design. Detailed engine changes have been proposed 

by Aerojet such as including the Gemini Stability Improvement 

Program (GEMSIP) injector on the Stage II engine and modifying 

the transtage engine propellant valve. These changes are also 

discussed in detail in subsequent sections. 

Incomplete crew safety and reliability studies may impose 

additional engine changes. The most significant Martin change 

anticipated will be the requirement for redundant hydraulic control 

systems for engine gimbaling. 

Several additional ground rules and assumptions were established to determine the scope and direction of the liquid systems 

comparison. They are: 

1) Only the YLR87-AJ-9 engine with the 15:1 expansion 

ratio was used on Stage I with the 7 seg-120 SRMs; 

2) Only the YLR87-AJ-9 engine with the 8:1 expansion 

ratio was used on Stage I with the 2 and 3 CS-156 

SRMs for performance studies; 
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VII-2 

3) Both the 15:1 and 8:1 expansion ratio engines were 

used for staging studies with the 156 in. SRMs; 

4) A nominal propellant temperature of 60°F was used 

for the WTR launch. A minimum of 45 and a maximum 

of 75°F was assumed; ' 

5) A maximum temperature of 90°F was assumed for staging 

NPSH studies on both the 8:1 and 15:1 expansion ratio 

engines for an ETR launch; 

6) Use of ground propellant temperature conditioning 

was not considered; 

7) Nominal propellant loading as defined in Section D 

was assumed; 

8) A minimum Stage I engine oxidizer NPSH of 35 ft for 

periods of approximately 4 sec was used for staging 

analysis. 

A. LIQUID ENGINE DATA 

A requirement was established at Martin contract go-ahead for 

Aerojet to supply data on the characteristics of the YLR87-AJ-9 

engine with the modified 15:1 expansion ratio thrust chamber and 

on the proposed modified YLR87-AJ-9 and AJ10-138 engines. The 

15:1 engine data were included with data on the present Titan III 

liquid engines and were used in the loads, performance, and 

staging studies. Table VII-1 summarizes and compares YLR87-AJ-9 

engine characteristics with 8:1 and 15:1 expansion ratio thrust 

chambers. The plume profile for the improved 15:1 engine is 

shown in Fig. VII-1. Table VII-2 summarizes the YLR91-AJ-9 engine characteristics, and Table VII-3 summarizes the AJ10-138 

engine characteristics. 

Table VII-1 shows that the NPSH requirement for the 15:1 engine increases by 1 ft on both the fuel and oxidizer pumps, which 

results from a shift in the pump operating points. This number 

represents the minimum steady-state run condition. 
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VII-3 

Ta ble VII-I StageIEngine Characteristics 

2 CS-I56 or 3 CS-I 56 

with YL1187-AJ-9 

(e = 8:1) 

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co G 

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Gas Constant, 68.29 ft/GR. 

ChamberPressure, 

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Titan III/MOL Exhaust Plume at 220,000 ft 

0 0 0 0 0 0 0 0 

0 0 0 0 0 0 0 

c0 csi c0 ■Ct trs 

4.4 4.4 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VII-4 SSD-CR-65-206 

( •u7) arripvd autrta lvoitu 
Parameter Value 

Altitude Thrust (lb) 

Altitude Specific Impulse (sec) 

NPSH Requirement (ft) 

Oxidizer 

Fuel 

Engine Wet Weight (lb) 

Gimbaled Moment of Inertia (slug-ft 2) 

Overall Length (in.) 

External Exit Plane Diameter (in.) 

101 ,000* 

310 (nominal) 

30 

100 

1276 (nominal) 

334 (nominal) 

110.62 (max) 

68.5 (max) 

*Thrust Chamber Value: 

F

TCA 

sp - W - Autogenous Flow 

Total 

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VII-5 

SSD-CR-65-206 

Table VII-2 Stage II Engine Characteristics 

Table VII-3 Stage III Engine Characteristics 

Parameter Value 

Altitude Thrust (lb) 

Altitude Specific Impulse (sec) 

Engine Wet Weight (lb) 

Gimbaled Moment of Inertia (slug-ft2) 

Overall Length (in.) 

External Exit Plane Diameter (in.) 

8000 (nominal) 

305 (nominal) 

239.2 (max) 

16.25 (max) 

80.85 (max) 

47.5 (Max) 
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VII-6 

An analysis of staging conditions (Section C shows that for 

90°F oxidizer temperatures the oxidizer NPSH for the Stage I engine falls below the required NPSH value for either type of engine during the staging sequence. Aerojet was requested to review the engine pump characteristics to determine the feasibility 

of running the engine below the specified NPSH for brief periods. 

The Aerojet analysis indicates the engines can run at an NPSH of 

35 ft for approximately 4 sec, that should be sufficient to pass 

through the staging transient with no serious consequences to the 

engine. Additional testing would not be required to demonstrate 

satisfactory pump operation at the low NPSH with the 15:1 engine, 

since other development testing is anticipated. Engine and pump 

demonstration tests would be required if the 8:1 configuration is 

used. Additional solutions to the NPSH problem are described 

next. 

Engine Isp and thrust defined in Table VII-1 for the 15:1 engine are nominal for an engine calibrated for 60°F propellants 

and having standard inlet conditions* with flight corrections applied. The corresponding values for the 8:1 engine are nominal 

at flight conditions also. 

The 15:1 engine overall dimension in Table VII-1 is measured 

from the Aerojet/Martin mounting frame interface to the exit plane 

of the thrust chamber. It does not include the TCA cover. 

Figure VII-2 shows the 15:1 engine assembly. The thrust chamber gimbal attachment point and toe-out angle are identical to the 

8:1 engine. A 2-in. clearance between thrust chambers, when in 

the neutral position, is achieved when an identical mounting is 

used for increased expansion ratio chambers. A device for maintaining thrust chamber nozzle separation during Stage 0 boostflight loads should be provided. A link connecting the thrust 

chamber exit protective covers could be used. Figure VII-3 shows 

that when the thrust chambers are gimbaled hardover in the pitch 

plane, the chamber extends as much as 10 in. beyond the 10-ft dia 

of the core. 

The 15:1 engine start transient side loads into the actuators 

during altitude start have been estimated by Aerojet to be 21,000 

lb. Neither the start transient side force nor the aerodynamic 

loads resulting from the thrust chambers protruding beyond the core 

skin line represent a problem for the MOL mission. However, these 

two conditions could necessitate the redesign of the hydraulic 

actuators if future missions indicate the need for a core sea level 

launch. 

*Oxidizer pump inlet pressure is 75 psia, fuel pump inlet 

pressure is 35 psia, and a 172,...cition exists. 

lit,LASSlii 
Staee I Titan 

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SSD-CR..65-206 V11-7 
30.93 in. 

Gimbal Point 

Note: Engine length = 100.8 in. (from 

gimbal point), Maximum exit plane 

outside diameter = 63,0 in. 

— Stage I Engine in Null Position 

(Engine Offset = 2.0167 deg) 

 Stage I Engine in Hardover, 

Snubbed Position (Engine Offset ■ 

4.88 deg) 

Stage I Engine in Hardover, 

Snubbed, Overtravel Position 

(Engine Offset = 5.38 deg),_ 

2.0167 deg 

4.88 deg 

5.38 deg 

10.0 in, 

9.0 in, 

4 in, 

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VIIVehicle 

Centerline 

Vehicle Outside Skin Line 

Fig. VII-3 Titan III/HOL Stage I Engine Exit Plane/Vehicle Outside Skin Line Layout (13:1 Expansion Ratio) 
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1111-9 

B. LIQUID ROCKET ENGINE CHANGES 

Many liquid rocket engine (LRE) improvements have been proposed by Aerojet. These changes would provide (1) increased reliability and crew safety, (2) improved payload capability, (3) 

ability to meet Titan III/MOL environmental and staging requirements, and (4) a solution to manufacturing and maintainability 

problems. No changes have been considered for the Stage I 8:1 

engine, but redundant hydraulic control alternatives now being 

studied could result in changes to the engine. 

1. Stage I Engine 

Dynamically Stable Injector - This modification will incorporate experience gained from the GEMSIP chamber. The injector 

will include baffles that may be regeneratively or tip-injection 

cooled and will improve the injection pattern. The injector will 

provide a more stable combustion margin and should not be a highrisk development item since previous experience is being used. 

Gearbox high-Speed Shaft Redesign - This modification will 

remove the resonant and operational frequency of the high-speed 

shaft assembly from a high load amplification region and increase 

the load-carrying capacity of the bearing. These improvements 

will be accomplished by changing the bearings and their arrangement on the shaft and by changing the shaft diameter and overhang. 

This change and the associated testing should not be a large technical risk, since the associated changes are common mechanical 

design engineering problems. 

Improved Regeneratively Cooled Chamber - Since the thrust 

chamber design is being changed to incorporate an ablative skirt 

at an area ratio of 6:1 and extending to 15:1, it is desirable to 

redesign the tube bundle to achieve decreased fuel pressure drops 

and increased propellant temperature capability. The redesign 

will provide a propellant launch temperature capability of 90°F 

with an 'R.bo of 0,80.* Design of a new thrust chamber is completed, 

• and Aerojet's experience in development of regeneratively cooled 

chambers should result in little program technical risk in making 

this change. 

*R

bo is the ratio of the tube burnout heat flux that will 

occur under given run conditions to that burnout heat flux theoretically established by design characteristics. This ratio is 

an indicator of the probability of tube burnout when operating 

an actual engine with real components. 
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VII-10 SSD-CR-65-2W 

Redundant Engine Shutdown System (RESS) - The possibility exists of a single malfunction of the pressure sequencing valve (PSV) 

system during a command shutdown requirement. This malfunction 

could cause catastrophic results if the RESS is not provided. 

Presently, the engine is shut down by a signal to the PSV override solenoid valves. These valves in turn vent the PSV actuation pressure overboard and permit spring shuttling of the spool. 

This action closes the thrust chamber valves. 

A RESS is proposed that consists of a squib-actuated normally 

open shutoff valve in each subassembly. The valve would be located in the oxidizer bootstrap line upstream of the gas generator 

oxidizer cavitating venturi and would close on receipt of a signal 

(the same signal that goes to the PSV override solenoids). If the 

PSV override fails to operate, the squib valve stops oxidizer flow 

to the gas generator resulting in turbine inlet pressure decay and 

turbopump slow down. When fuel pump output pressure decays to 

approximately 300 psi, the spring closing feature in the fuel 

valve actuator will cause the main propellant valves to close completing the engine shutdown sequence in a known manner. The squib 

valve assembly weighs approximately 21/2 lb and should not present 

a risk, since it has been successfully used on the Stage II engine 

of Gemini. 

Imgroved Vaned Elbows - The propellant inlet vaned elbows have 

historically shown distortion and cracking following engine shutdown. Although no structural failures occurred, the potential 

loss of a vane or portion of a vane could result in downstream 

blockage of a combustion chamber tube, injector passage, or orifice. It is proposed that the elbows be redesigned to eliminate 

vane distortion and cracking. This redesign and testing would 

provide added confidence in the engine system and should not cause 

an undesirable technical program risk. 

Altitude Nozzle Extension - This modification will provide 

added thrust and specific impulse performance for altitude operations of the engine. A glass fabric-wrapped honeycomb structure 

with a silica phenolic liner is proposed similar to the Titan III 

Stage II nozzle extension. Design analysis is completed and the 

redesigned nozzle extension is shown in Fig. VII-4. Experience 

with ablative skirt testing and development at Aerojet indicates 

a negligible program risk from redesign of this item. 
A ll dimensions are in in 

cU 4.1 

O 

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RELEASE 1 JULY 2015 

vu-n 

00 

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> 4,4 

o .c 

O4 1-1 

CO 

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15:1 Stage I Ek;ine Ab lative Skirt 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VII-12 

Optimized Fuel Pump - A decrease in the regeneratively cooled 

tube jacket pressure drop in conjunction with a trimmed fuel pump 

impeller and a slight increase in turbine speed can provide additional thrust to the engine with no increase in turbine horsepower. Thus, an impeller trim has been proposed with the required 

impeller discharge vane angle change. The impeller housing would 

not be modified. Figure VII-5 shows the proposed impeller trim, 

Some technical risk is involved with this modification, but theoretical analysis indicates only small changes will be required. 

Similar development has not been conducted by Aerojet, but because 

of small design variation, the program risk should be small. 

Subcomponent Modifications - Other small changes have been 

proposed, e.g., incorporation of more corrosion-resistant materials 

in the gas cooler and superheater, a start cartridge temperature 

compensating nozzle, mechanically locked turbine blades, etc, which 

should be given further consideration. These items could reduce 

maintenance costs and improve reliability. The technical risk 

associated with incorporating Most of these items should be small. 

2. State II Engine 

Similar improvement items have been suggested for the Stage II 

engine but are limited to the following: 

1) Incorporation.of GEMSIP injector; 

2) Incorporation of augmented engine improvement program (AEIP) items; 

3) Improved combustion chamber; 

4) Redundant engine shutdown system; 

5) Subcomponent improvements. 

These changes have been made on the Gemini launch vehicle except for the AEIP improvements, improved combustion chamber, and 

the subcomponent improvements. The AEIP items were evaluated 

early in the Titan III program and some were developed, thus, in 

making these changes, the program technical risk is minimal. The 

combustion chamber change should require only a small variation 

in tube size with a resulting improvement in high-temperature 

propellant operation. Little program risk is anticipated. Subcomponent improvements are characteristic of the Stage I subcomponent improvements and are considered to be of equally low risk. 
VII-13 

SSD-CR-65-206 

Pump Housing --- 

11.85 

in. dia 

10 77 

in dia 

Impeller Tip Profile 

10.48 in. dia 

Modify Impeller 

Per This Dimension 

Impeller 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

Fig. VII-5 Proposed Improved Fuel Pump 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VII-14 

Stage III Engine 

Studies conducted by the transtage engine manufacturer have 

indicated a need for only one change. A modification of the propellant thrust chamber control valve to improve engine start and 

shutdown transient characteristics is, desirable. A single pilot 

valve to control both engines has been proposed using a single 

hydraulic actuation pressure source. Reduction in system dynamic 

interaction may be accomplished with this independent pressure 

source. An electrically actuated system also was considered, but 

a hydraulic system appears to be most desirable. Since the transtage engine characteristics are acceptable with its current propellant control valve, little program risk will be associated with 

further testing and development of this item. 

It may be desirable to provide a RESS on Stage III, as well, 

as Stages I and II, for redundancy. The addition of a RESS 

should be studied further. 

Some of the engine modifications are not directly related to 

performance improvement but to crew safety or reliability. All 

modifications must be evaluated to arrive at a valid overall assessment of engine changes. 

C. STAGE I/SRM STAGING NPSH STUDY 

1. Study Criteria 

The Stage I minimum interface NPSH required is tabulated 

below. 

Engine Expansion Ratio Fuel Oxidizer 

8:1 43 ft 44 ft 

15:1 44 ft 45 ft 

The two major differences between Titan III and Titan III/MOL 

that affect NPSH are decreased acceleration during staging because of larger solids and increased NPSH required by the 15:1 

engine. The increase in NPSH requirements is relatively insignificant. The decrease in acceleration results in a decrease in the 

Stage I oxidizer NPSH of as much as 15 ft for the 3 CS-156 configuration. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VII-15 ' 

Because of the placement of the fuel tank with respect to the 

engine, acceleration effects on NPSH are less significant than in 

the case of the oxidizer tank. Also, more margin exists between 

required and supplied NPSH, and therefore, the fuel tank can meet 

the required NPSH criteria for all Titan III/MOL vehicle configurations with no modifications. 

2. NPSH Improvement Items 

The following changes were considered to provide increased 

Stage I oxidizer NPSH. 

Reduction in the Sta)'e I EnT ne Oxidizer 8xidizer0PSB Re ed - A 

study of existing pump, TPA, and engine data gathered by Aerojet 

includes significant test history of oxidizer pump operation below minimum NPSH. These data are sufficient to conclude that 

TPA component reliability and performance will not be significantly degraded at 35 ft oxidizer NPSH or above for short periods. 

Engine tests at low oxidizer NPSH do not indicate any adverse effects on system reliability or performance. However, additional 

demonstration tests under Titan III/MOL conditions will be required. 

Increased Oxidizer Propellant Tank Lockup Pressure - A study 

performed by the Martin Company's Stress Group shows the maximum 

tank pressure before liftoff can be increased 7 psi, thereby increasing NPSH at the critical staging point by approximately 5.0 

ft. This study was based on a maximum acceleration during solid 

motor burn 'of 3.2 g. This increase requires only a change in the 

launch limit pressure switch settings. 

Increased Minimum Oxidiz Ulla e - A small increase in ullage 

accompanied by a shift in mixture ratio could result in approximately 1-ft increase in NPSH at the critical staging point, but 

further ullage increases would result in a significant payload 

penalty. 

Oxidizer Autogenous Gas Flow Box Change - This would require 

'a redesign of the Stage I superheaters to gain significant NPSH 

increases resulting in payload penalty due to increased residual 

gas weight. 

Reduction in the Maximum Propellant Temperature Limit - This 

reduction would slightly reduce the launch-on-time probability 

at launch sites requiring maximum propellant temperatures of 75°F. 

Propellant conditioning would be required for launch sites requiring maximum propellant temperatures of 90°F. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VII-16 

Gas Injection System to Pressurize the Oxidizer Tank during 

Staging - Studies showed that this would be technically feasible, 

but the development required, cost, weight, operational problems, 

and the possibility of accidental pressurization resulting in 

pressures approaching structural limits make this change practical only if all other solutions fail. 

3 S d zer NPSH Stud Results and Recommended ol 

A study has been performed to determine the Stage I oxidizer 

NPSH supplied for the different Titan III/MOL vehicle configurations. This study was limited to three basic SRM configurations 

since the acceleration effect on NPSH varies less than 1%, with 

and without transtage. Staging acceleration data for the three 

configurations were studied.* The autogenous pressurization system 

7094 machine program was used to generate minimum oxidizer tank 

top pressure curves. The following assumptions were input to the 

program and were based on Titan III flight data. ' 

1) Autogenous gas was supplied at the minimum point on 

the Titan III Martin/Aerojet interface flow box. This 

point was reached at a steady-state condition at 87FS1 

+ 1.3 sec for the gas flow and 10 sec for enthalpy; 

2) Steady-state propellant flow was reached at 87FS1+ 

1.5 sec; 

3) The minimum tank pressure at 87FS1 was 34.7 psia. 

This included the minimum lockup pressure less the 

effect of pressure decrease due to prevalve opening 

and tank pressure decay during Stage 0 burn; 

4) Minimum pressure at 87FS1 was increased to 41.7 psia 

to include the allowable 7-psi increase in lockup 

pressure; 

5) Minimum ullage of 85 cu ft was present at 87FS1. 

The acceleration data from TM 5141/31-65-19 and tank top pressure 

data were used to determine NPSH supplied for the conditions tabulated on the following page. 

*D. Bressler, S. Bonson, W. Livesey: MOL Compatibility Solid 

Rocket Motor Staging Analysis Report. TM 5141/31-65-19. Martin 

Company, Denver, Colorado, 1965. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VII-17 

Configuration 

Engine Expansion 

Ratios 

P'ropell'ant 

Temperature 

(°F) 

Propellant 

Flow Rate 

(cu ft/sec) 

Lockup 

Pressure 

(psia) 

7 seg-120 15:1 75 12.50 37 

7 seg-120 15:1 90 12.55 37 

2 CS-156 8:1 75 12.05 37 

2 CS-156 8:1 90 12.10 37 

3 CS-156 8:1 75 12.05 37 

3 CS-156 8:1 90 12.10 37 

3 CS-156 8:1 90 12.10 44 

Figures VII-6 thru VII-8 and VII-10 show that a reduction in 

required NPSH satisfies all configurations at a maximum propellant 

temperature of 75°F and the 7 seg-120 SRM configuration with 90°F 

propellants. Figure VII-9 shows the 2 CS-156 SRM configuration 

with 90°F propellants, which drops slightly below the 35-ft NPSHrequired curve. Figure VII-11 shows the 3 CS-156 SRMs with 90°F 

propellants. NPSH supplied falls significantly below the 35 ft 

required. Figure VII-12 shows this same configuration with increased lockup pressure. The figure reflects that minimum NPSH 

for the worst case configuration can he met by decreasing the required minimum NPSH and increasing lockup pressure. 

4. Conclusion 

The only three items that significantly decrease the difference 

between the supplied and required NPSH are decreased required NPSH, 

increased lockup pressure, and the installation of a gas injector 

system. None of these represents appreciable technical risk. The 

first two do not require major redesign, therefore, the recommended 

Solution is to decrease NPSH requirements for the 7 seg-120 SRMs 

and decrease NPSH and increase lockup pressure if the 2 and 3 CS156 SRM configurations are selected. 
80 

NPS71 at 87FS 

Note: 1. Expansion ratio - 15:1. 

2. Minimum tank lockup 

pressure - 37.0 psia. 

0.1.1.• //femme ...maw •••■■•■■ Hem. 

40 

al• ■•••••••■ 

20 

p sed ni 

4 5 6 7 8 9 

Time (sec) 

Fig. VII-7 Titan III/MOL Stage I Oxidizer Minimum NP$H during SPM Staging 

(7 aeg-120; 90"F Propellant Temperature) 

11 12 

4 6 7 

Time (sec) 

9 0 11 12 

7 

100 

a 

40 

20 

1■•■■1011. ••••■49 10001•111 VaTIMM 

Expansion ratio = 1517. 

Minimum tank lockup 

pressure = 37.0 psia. 

4 

, I Required - T -77 

Proposed Minimum NPSH 

80 

t: 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

Vil -18 

PSII Required 

NPS 

*WM/. rano.,0•1•41.5, •■•■, 

10 
0 0 0 

op 

0 0 

(n) ItScIN 

0 

co 

0 0 0 

tNi 

(n) Hum 

1 i 

1 ill 

Expansiot 

Minimumi 

pressure 

II , iat87FS I 1_ 1 I 

I 1 

II • il bill 

NI 

Est 

• 

 NP: 

. 

iii 

Note: 1. Expansionratio =8:1. 

2. Minimumtank lockup 

pressure= 37.0 psia. 

1 

li 

H fl ii NPSR Requ 

.... 

44 

N4 

CO 

4-4 

v-4 

4-4 

OD 

0 

0 DD 

4ti 4 

GO 

rn 0 

■•••• 

W 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VII-19 

SSD-CR-65-206 
...•■■■••■ al.“•■••■ 

8 9 10 11 12 2 3 4 6 7 

Time (sec) 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VII-20 

80 

Note: 1. Expansion ratio = 8:1. 

2. Minimum tank lockup 

pressure = 37.0 palm,. 

10 NPS11 at 87FS 

1- 

Minimum NPSH Required 

'"--1 

Proposed Mini num 40 

20 

100 NPSH at 87 

60 

40 

20 

U 

80 

Note: i, Expansion ratio = 8 1. 

2. Minimum tank lockup 

pressure = 37.0 pale. 

Fig. VII-10 Titan III/MOL Stage I Oxidizer Minimum NPSH during SRM Staging 

(3 CS-156; 75.F Propellant Temperature) 

0 2 3 4 5 6 7 a 9 10 11 12 

Time (sec) 

Fig. VII-11 Titan III/MOL Stage I Oxidizer Minimum NPSH during SRM Staging 

(3 CS-156; 90.F Propellant Temperature) 
a 

00 

CO 4.) 

Crl P. 

0 

cn 0 

00 

Y4 CA 

1:3 

CP, 

Ci) 

p4-4 

g o) 

00 

Cd 

}-1 

Cn 

1.4 5 

4) 41 

Cs1 H 

't1 

0-4 

X al 

1-1 41 

a) a. 

0 0) 0 

004-4 

cd P4 

0 

C.0 

0 0•' 

cn 

!.! J C"n O 

00 

0.1 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

V I - 21 

- A . 

mmNPSH Req

uired! 1 

— --r — — I -r 4-—l— 

ed MinimumNPSH 

1, Exp

ansion r

atio = E

2. Minimum ta

nk lockuf 

Pressure = 44.0 psi 

Minim! 

Prop

os 

al 

al 

N

PSH at 87FS, 

I I 

Ell 

<NI 

C.) 00 

(33) liSd61 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VII-22 

D. PROPELLANT LOADING INVENTORY 

A detailed breakdown of the propellant loading inventory is 

given in Tables VII-4 and VII-5. Two transtage loads are given; 

Table VII-4 reflects the off-loaded-case and Table VII-5 the 

fully loaded case. The calculation of the loading is basically 

the same as that used in the Titan III program. The following 

criteria were used in the analysis. 

1) The engine nominal mixture ratios were selected to 

maximize the propellant loads; 

2) Since Stage I and II loads are a function of propellant temperature, a nominal temperature range had to 

be selected. The ranges of propellant temperature 

blocks for this study are 40 to 55, 50 to 65, and 

60 to 75°F. Therefore, for this study the nominal 

propellant load is based on the 50 to 65°F range; 

3) All other figures used represent the latest information available from the Titan III program. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VII-23 

Table 111-4 liquid Propellant in,ntory T.ft,0 551i5101 (Tra4at,,ge (9 

---- -- - 

Item _ Stage 1 „ 

5010 61 

1.806 

Slate III 

1. Engine Number 

2. Average infilght Mixture Ratio 1.91 

3. Propellant Temperature CO K to 65 

' 

----- 45 64 2f, 

l%'Eld 

---------- 

4, Propellant Density 

__...---- 

1,571.87 

1,864.37 

---- 

430,45 

482.06 

42.1909 

" 

---- 

Oxid 

--- 

Fuel 

5. Maximum Loedable Volume (Cu ft) Oxld 

6. Nominal Propellent Loaded (16) Fuel. 

Tote-1257,390 ,102 

88,795 

168,595 

24,316 

43,593 

5,381 

Oxld 10,721 

7. Propellant Expended before 

Liftoff (16) . • 

Total 6 , 

Fuel, • 

6 

0 

2 

Oxid 

- a, Engine Bleed 

6. Engine Leakage 

c. Start Consumption, 87FS, 

. to TOPS 

d. Holddown Conaumption, TCPS 

to Liftoff 

Fuel 6 

0 

2 0 

0 - ,-- 

°mid 0 

1,,e1 , 0 

N/A 

o 

Oxid 0 . 0 

Fuel N/A N/A 

Oxid 

Fuel N/A 

233,384 

N/A N/A 

Oxid 

Total 

8. Propellant Aboard at Liftoff 

67,907 16,102 

Fuel 

Oxid 

88,789 

168,995 

24,314 

43,593 

5,3d, 

10,721 

9. Propellant Expended during Previoue Stage Operation 

Total 29 15 

28 

1 

15 

Oxid 0 

a. Engine Bleed, Stage 0 Operalion 

b. Engine Leakage, Stage 0 

Operation 

c. Engine Bleed, Stage I Operalion 

d. Engine Leakage, Stage I 

Operation 

Fuel. 28 

0 

6 

Oxid 0 

Eel 1 

, 1 

0 

Oxid 0 

Fuel N/A 9 

Oxid 0 

Fuel N/A 

0 

0 I Oxid 

10. Propellent Aboard at 87FS1 Total 

Fuel 

257,555 

,z, 3 

168,594 

67,092 

- 

6,3.2 

24,299 

43,593 

5,.16: -

Oni? '

10,121 

11. Engine Leakage during Stage 

Operation 

Total .12 0 

FueL 11 

1 

7 

67,480 

24,142 

43,338 

0 

16,088 

5,380 

10,708 

Oxid 

12, Total Available Ueable Propellent 

Total 

Fuel . 

Oxid lbiD4 

a. Start Consumption 

b. Steady-State Consumption. 

e. Shutdown Consumption 

J. Talloff (before Staging) 

e. Fuel Bias 

Fuel 55 

138 

2 

Oxid 

Fuel 94/ 5}( 

, 

23,912 

43 

5,350 

Oxid 10,700 

Fuel 5 

94 

 1 

Oxid 5 

Fuel NIA 

, 

24 

N/A 

O xid 

Fuel -- - 285 96 25- 

13. Total Nonusable Propellant 

Total 1,135 404 14 i .. ....... 

Pool 446 

689 

150 

254 13 Oxid 

a. Propellant Vapor Retained 

6, Trapped above Interface 

c, Trapped below Interface 

Fuel 99 

515 

613 

193 

3 

13 Ogle! 

Fuel 

Oxid 

50 

0 

61 

23 

0 

0 

-11.124-..- gl 

174 

29 

38 

0 

Oxid 0 

14. Mean Outage Total 548 215 50 

15, Nominal Propellant Consumed 

during Stage Operation (Item 

12 - 14) 

Total 255,660 67,265 16,038 

a. Nominal Steady-State Propellent (Item 126 - 14 + 12e) 

6. Transient Propellants 

Total 255.082 66,875 16,025 

Total 

--- 

278 390 13 

16. Pressurization System Inert Gaa (ZTI 

4 2

9 

60 

Oxid 15 

Spheres N/A N/A 45 

17. Maximum Outage 2,100 675 161 
10 65.206 

Table VI/-5 Liquid Propelle,t Invo (Trnalaft:04 Fully Loaded) 

Item Stage I Stage II Stage III 

1. Eng e Number 

2. Average In-Flight Mixture Retie 

Oxid 

rd 

3. Propellant Temperarure (°F) 

10, 

45 to 75 

4. Propellant Density 

Vial 137,33 

5. 172.65 Maximu m Loadable Volume (cu ft) d 

6. Nominal Propellant Loaded (lb) 

To 23 081 

Fuel 

Oxid 

7, Propellant Expended Before 

Liftoff (16) 

Tot al 

KIM 

111111101 1Min e 

23,061 

Oxid 

a. Engine Bleed 

b. Engine Leakage 

c. Start Consumption, NMIto 

TCPS 

d. Holddown Conaumption, TIPS to 

Liftoff 

FHeI 

Fu 

uel 

Oaid 

Fuel 

Oxid 

8. Propellant Aboard at Liftoff 

Total 

7,11 

IIMM 

13,367 

0 

aIIIIIIIIIIIIIMIIII 

MEI 

IIIMIIIIIMIIII iw__

23 L..0 81 

7,714 

ME 15,367 

0 

Oxid 

9. Propellant Expended during Pre 

ViOUIP Stage Operation 

otel 

Fuel 

U 

a. Engine Bleed, Stage 0 Operation 

b. Engine Leakage, Stage 0 

Operation 

c. Engine Bleed, Stage I Operation 

d. Engine Leakage, Stage I 

Operation 

Putt 

Oxid 

Fuel 

10, Propellant Aboard at iliFS 1 Total 

u 

Oxid 

11. Engine Leakage during Stage Operation 

Total 

u id MN 

12. Total Available Doable Propelbust 

To 23 067 

Fuel 7,713 

L 15,354 

a. Start Consumption 

b. Steady-State Consumption 

c. Shutdown Consumption 

d. Tail off (before Staging) 

e. Fuel Biro 

F13

o 2 

•

e 

 15,346 

7,673 

l.Fxid 

Furl 

1 

35 Fuel 

allM 

13. Total Nonuaable Propellant 

a. Propellant Vapor Retained 

b. Trapped above Interface 

c. Trapped below Interface 

duel 

Tot el 

x d 

t 0 

Oxid 

0 1d 

Puel 

Mal 

14. Mean Outage 

15. Nominal Propellant Consumed during 

Stage Operation (Item 12 - 14) 

a. Nominal Steady State Propellent (Item 126 - 14 + 12e) 

b. Transient Propellants 

Total 

22,995 

22,982 

13 

16. Pressurization System Inert Gee 

total 

7 

8 

OS 

6'' xid 

Spherea 

17. Maximum Outage 231 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VII-24 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VIII. PAYLOAD LENGTH 

This chapter presents the results of studies to determine 

the effect on allowable payload length of reduced launch probabilities and structural redesign of the core structure. The 

results of the more important loads analyses are also discussed. 

The primary loads study was to determine the maximum payload 

length that could be carried by the six reference configurations 

(see Chapter II) based on 99% launch probability and no redesign 

to increase the basic longitudinal strength of the core. Secondly, the reduced launch probability for the specified payload 

lengths was to be determined, again based on no redesign to 

increase the core longitudinal strength. Thirdly, the core 

weight penalties associated with structural modifications 

required to attain 99% launch capability were to be determined. 

PAYLOAD DESCRIPTION 

The MOL payload configuration consists of the Gemini plus 

laboratory module. To obtain parametric data, it was necessary 

to select two additional payload lengths for each of the three 

solid rocket motor (SRM) configurations, Table VIII-1 lists 

the specified lengths plus two additional lengths for each configuration. The additional lengths were selected to give a 

reasonable length spread for parametric studies. The center of 

gravity (cg) locations were provided by the customer for the 

specified lengths. Weight distributions for the other payload 

lengths were assumed such that the resultant cg would stay at 

the same percentage of overall payload length. The payload 

weights that were used for each SRI configuration are also 

listed in Table VIII-l. 
S Payload Weight (lb Length (ft) X (ft 

54.5* 29.6 

7 seg-120 28,000 65.0 35.2 

75.0 40.6 

50.0 27.1 

2 CS-I56 33,000 58.5* 31.7 

65.0 35.1 

50.0 27.1 

3 CS-156 42,000 61.0* 33.1 

72.0 39.2 

*Specified p ayload lengths. 

Length 

Center of 

Gravity 

TIII/140L 

Interface 

10.5-ft Diameter 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VIII-2 

Table VIII-1 Nominal Payload Lengths Used for Loads Analyses 

(With-Transtage Configuration) 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VIII-3 

B. RIGID BODY TRAJECTORY RESULTS 

A rigid body airload study was performed on each of the 

three with-transtage vehicle configurations. The following 

baselines were established for the analysis: 

1) Rotating, oblate earth; 

2) Rigid body dynamics; 

3) Nonlinear aerodynamics; 

4) Load relief autopilot; 

5) Launch azimuth of 182 deg from WTR; 

6) Standard atmosphere of 1962; 

7) Design wind - Taken from Meteorological Note 2.* 

In all cases, the wind azimuth search was conducted using a 

wind shear peak altitude of 29,600 ft. This is the lowest altitude at which the maximum wind magnitude and wind shear values 

occur for the wind criteria used. After the critical szimuth 

was determined, i.e., the azimuth that produces the largest 

value of the airload parameter qCO, various wind shear peak 

altitudes were introduced over the altitude range of 11,000 to 

35,000 ft along the critical azimuth. In this manner, the azimuth and shear peak altitude combination that produces the largest value of qa8 was determined. Table presents a summary of the most significant trajectory results for the critical 

flight conditions. 

Some trajectory analysis was conducted using reduced winds. 

Figure VIII-1 presents a plot of quo vs the percentage of maximum wind for the 7 seg-120 configuration. Since the wind is a 

pure'side wind, clop varies directly with p. These data were 

used in conjunction with wind load analyses for full winds to 

obtain design data for reduced wind bending moments. 

Jerold Bidwell: Atmos•heric and Wind Desi n Criteria or 

PMR, Meteorological Note 2, is r in Company, Denver, Colorado, 

7 August 1964, 
1.1 

"0 

0 C.) 

61) 4,4 

tr. 

4.4 

J.J 

4.4 

1-4 

0 

U •■-■ 

(1.1 

4-4 

S.1 

1.1 

V 

4-1 

E-+ 

H H H 

V 

E-+ 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VIII-4 

A irload Angle of 

Sideslip, e Indicator, qdroft 

deg ( ) (1b-degift2) 

(--) 

-- 

Lri 

-.1 

-..1 

tn 

....7 

\C) ,t) 

.1' 

aN 

an 

.---1 

0 CO 

st ■0 

in 

5.80 4901 

5.78 4900 

5.78 4900 

tn 

1--. 

-4' in 

..I 

r 

V) 

un 

co 

0 

Ln 

N. 

it) 

Cr' 

in 

N 

In 

Dynamic 

Pressure, q (Ibift2) 

g 

00 

,--1 

c0 

CO 

ON 

r-- 

c0 

1.,-1 

-Ic0 

N 

-4- 

00 

N 

...,1. 

cc 

co 

...1- 

o \ 

co 

....t 

a\ 

N 

....1- 

0' 

= 

tif 0 Z Z 

cc ■-i 

c0 

1--+ 

N. N 

,--4 

N. 

ri 

N.; lf) 

,-.1 

WindShear 

Peak 

(ft) 

32,000 

32,000 

32,000 

0 

01 N 

0 0 vP 

a 

ON Cs4 

0 0 

4, 

a 

at N 

009'6Z 

009'6Z 

009'6Z 

Timeof Flight 

(sec) 

0 

0 tn in in 

at 

01 

H 

0' -3- 

.—t 

a., in 

H 

kr) 

st 

0 

kr, -4' 

..= 

"V = 00 

c s <1.1 

• r1 

•,..1 '0 3 ty ...■ < 

N N. 

CN 

N N N 

N N. N 

N N N 

N 

N. N 

N 

N N 

trb 

4 N 

V) 

...,1" N 

un 

....t N 

I 

Payload 

SRM Length 

Configuration (ft) 

1-n 

.4' LP l 

0 

in 

..0 

CD 

N 

a-1 

1 

00 V 

co 

N 

0 t-41 N. 

0 

0 

in 

in 

00 LP 

k.0 

Lil 

t 

cn 

C.) 

N 

Jr) 

v.-4 

0 

0 

1 

0 

t--1 %JD 

\i) 

Lrl 

i 

cn C.) 

rn 

0 

N ' 

N. 
VIII-5 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

0 

0 

v.) 

0 

0 

( 0ap.ca) gob cacnimput speoTaw 

‘Z 

000 d 

'cIT IsepTs go 

a

i2uv 

(290 70 10v431/ go alituif 

Maximum WTR Wind Speedand Wind Shear (Z) 

Fig. VIII-1 Airload Indicator vs Percentage of 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VIII-6 

C. LOADS ANALYSIS ASSUMPTIONS 

Early in the 60-day study, certain ground rules and analysis 

assumptions were agreed upon with Aerospace to facilitate arriving 

at common answers. The more significant of these were: 

1) Design wind criteria will be based on Meteorological 

Note 2.

, 

 It was then later agreed that the launch 

probability associated with reduced winds would be 

based on data from the National Weather Records Centert 

for the worst month, This relationship between percentage of wind and launch probability is shown in 

Fig, VIII-2. 

2) A buffet analysis will be conducted; however, dispersion and gust effects will be estimated based on previous studies. Dispersions used were 28% of pitch moments and 20% of yaw moments. Gust effects were estimated to vary from 25 to 35% along the vehicle length. 

3) For switchover analysis, a more realistic method of 

loads combination than that used for nonmalfunction 

conditions will be used together with an ultimate 

safety factor of 1.25. This was decided primarily 

due to the high improbability of a malfunction occurring at exactly the same time as the maximum wind 

shear spike. 

4) The ultimate safety factors used in other loads analyses will be consistent with those used for Titan 

IIIC design. 

5) For a given payload length, the changes in core bending moments due to inertia relief differences are 

small and will not be accounted for in the 60-day 

study, 

6) The existing strength of the core will not be corrected for changes in local aerodynamic collapsing 

pressures and elevated temperature effects, This was 

decided since the core temperatures and local collapsing pressures for the three SRM configurations 

are not much different from each other or from Titan 

IIIC values, 

bid. 

tWinds Aloft Summary and Parameters, Pt, Arguello, California 

No. 4647, National Weather Records Center, 2 December 1963. 
h Probability 

80 

70 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VIII-7 

100 

90 

60 

50 

50 60 70 80 90 

Design Wind (%) 

Fig. VIII-2 Launch Probability vs Design Wind 

100 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VIII-8 

D, MAXIMUM AIRLOAD AND LENGTH ANALYSIS RESULTS 

1. Core Loads 

The bending moment and axial loads corresponding to the critical flight conditions of Table. VIII-2 are shown in Fig. VIII-3 

thru VIII-5 in the form of PEquivaent curves for each of the 

three SRM configurations. The present core capability is also 

shown in each of the figures. All loads shown are based on 

100% of design wind. Since all loads analyses were conducted 

for a configuration including the transtage with the payload 

interface at Station 77, it was necessary to convert the nominal 

payload lengths to actual lengths for both the with- and withouttranstage configurations. These corrected lengths and the appropriate interface stations are listed in each of the figures. 

The axial loads shown include transtage propellant weight. 

The bending moments for both the 7 seg-120 and 3 CS-156 

configurations peak at Station 250, since that is the location 

of the SRM-to-core forward structural tie. In a like manner, 

the moment for the 2 CS-156 configuration peaks at the same 

station as that for the Titan IIIC 5-segment SRM configuration. 

In general, the actual PEquivalent for the long SRM configurations 

exceed the present allowable at the forward end of the core and 

then become less critical aft of that location, while the 2 CS156 is generally more critical over a longer portion of the length 

of the core. 

Since the core structure forward of Station 296.6 is being 

redesigned for the without-transtage configuration, that station 

is used as the forward limit of the core allowable for each of 

the without-transtage configurations. 

2. Allowable Lengths 

The information from Fig. VIII-3 thru VIII-5 was used to 

plot,the curves in Fig. VIII-6. Using the three length data 

points for each of the six configurations, plots of the ratio 

of actual to allowable loads at the critical station are presented. The intersection of the six curves with the ordinate 

value of 1.0 represents the allowable payload length that can 

be carried based on 99% launch probability and no basic core 

longitudinal strength increase. These allowable lengths plus 

the specified lengths are tabulated in Fig. VIII-6. Corrected 

payload lengths were used and the axial load for the withouttranstage configurations took into account the deletion of the 

transtage propellant weight. 

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III-5 

NRO APPROVED FOR 

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f1)-CR-65-206 VIII-11 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 

V111-12 

Configuration Station 

Allowable 

Length (f 

Specified 

Length (ft)., 

7 seg-120 with Transtage 151,6 57.8 54.5 

7 seg-120 without Transtage 296.6 74.7 74.5 

2 CS-156 with Transtage 320.0 51.3 58.5 

2 CS-156 without Transtage 320.0 67.8 78.5 

3 CS-156 with Transtage 151.6 53.0 61.0 

3 CS-156 without Transtage 296.6 71.2 81.0 

Fig. V111-6 Allowable Payload Lengths, 99% Launch Probability 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

3. Reduced Launch Probability 

Information from Fig. VIII-1 and VIII-2 plus the basic airload analysis was combined to present the curves shown in Fig, 

VIII-7. The second study objective results, launch probabilities 

for the specified payload lengths based on no redesign to achieve 

increased core strength, are tabulated in the figure. Also shown 

are the launch probabilities for the maximum length payloads investigated, 70 ft for the with-transtage configurations and 82 ft 

for the without-transtage configurations. 

uctural Redesign 

The results for the third objective of the payload length 

study are shown in Table VIII-3. Listed are the delta core 

structural weights associated with a redesign to permit 99% 

launch probability for each of the specified and the maximum 

payload lengths. The 7 seg-120 totals for the specified lengths 

are zero since that configuration already has 99% launch probability with the present strength. In general, the weight penalties and the extent of redesign are more severe for the 2 CS-156 

than for the 3 CS-156 configuration. 

Summary and Conclusions 

In previous paragraphs the possibilities of reducing the 

launch probability and redesigning the core structure to achieve 

greater payload length capability have been discussed. Table 

VIII-4 summarizes that information and includes two other possibilities of potential payload growth. These are a reduction in 

the design ultimate safety factor from 1.4 to 1.25 for airload 

bending moments and also a reduction in the design value of q43. 

The reduction in design ultimate safety factor will necessarily 

include an associated increase in the number of structural 

failures with less than 3 sec of warning time. Preliminary calculations indicate an increase of about 50. A reduction in qc3 

could be achieved by changing the flight trajectory somewhat and 

also by redesigning the solid rocket propellant grain. Such 

studies were beyond the scope of the 60-day study. Note that a 

significant payload length increase could be achieved by a combination of several of the methods shown. 
Launch Probabilit,ie (%) 

'Specified 

Configuration Length 

7 seg-120 with Transtage 

7 seg-120 without, Transtage 

2 CS-156 with Transtage 

2 CS-156 without Transtage 

3 CS-156 with. Transtage 

3 CS-156'without Transtage 

>99 

99 

94.5 

89.5 

88 

83.5 

*70 ft for with-transtage configurations; 82 

without-transtage configurations. 

Paylo

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100 

90 

80 

70 

60 

50 

40 

0 60 70 80 90 100 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

V111-14 

ax. Length* 

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Launch Probability (%) 

Fig. V111-7 Allowable Payload Length vs Launch Probability 
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III-4 

NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VIII-16 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

VIII-17 

E. OTHER LOADS ANALYSES 

Several loads analyses were conducted during the 60-day study. 

The more important of these and their results are; 

1) A launch analysis considering differential thrust 

buildup between the two solids resulted in loads 

less critical than those at staging; 

2) The maximum airload conditions were the primary 

contributors to the payload length studies just 

described. The Gemini/laboratory module bending 

moments for the critical flight condition for the 

specified payload lengths are shown in Table VIII-5; 

3) Step 0 burnout was not quite as critical as the unsymmetrical tailoff condition at the time of Stage I 

engine start. The most important result of this 

analysis, the Stage I longeron load, is discussed in 

Chapter V; 

4) The switchover loads analysis results indicate adequate structural margin to permit approximately a 

100-msec time delay in switchover; 

5) Thrust termination results cannot be taken as conclusive at this time since thrust termination curves 

for the solid configurations being studied were not 

available. However, the requirements that have been 

levied on the SRM manufacturers and the inputs to our 

loads analysis are such that only local redesign is 

required to the core/SRM attach hardware. Substantiation that additional redesign is not required 

must await receipt of actual thrust-termination 

curves for the selected SRM configuration. 
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NRO APPROVED FOR 

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V111-18 

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NRO APPROVED FOR 

RELEASE 1 JULY 2015 

IX-1 

IX. RELIABILITY AND CRE4n-5,AFETY 

The objectives of the reliability and crew safety program 

study were to (1) analyze the six performance configurations 

with respect to their crew safety and mission success differences to provide a basis for performance modification and selection and, (2) analyze a Titan IIIC for an updated mission success 

evaluation. SSD-CR-65-203* includes a comprehensive summary of 

the reliability and crew safety program. 

A. GROUND RULES 

The following ground rules were established: 

1) The analyses would be based on predictions for the 

first Titan IIIC/MOL booster launch (1968); 

2) The analyses would be oriented to achieved predictions and would include the expected degradation due 

to manufacturing, assembly, and test errors; 

3) Mission abort was defined as any failure that prevents 

achievement of the primary orbital mission; 

4) Only single malfunctions leading to mission abort 

would be considered except for the thrust vector 

control (TVC) subsystem, which will include multiple 

malfunctions. Abort due to multiple malfunctions 

are considered to have little impact on overall mission aborts; 

5) All equipment except interconnections and cabling 

would be analyzed; 

6) The analyses would be based on mission time profiles, 

which will include the different SRM burn times and, 

a 310-sec first burn and an 8-sec second burn for 

with-transtage configurations; 

*Preliminary Failure Mode and Effects. SSD-CR-65-203. Martin 

Company, Denver, Colorado, 27 August 1965. 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

IX-2 

7) The analyses, would include Inputs from each of the 

participating associate contractors; 

8) The Gemini inertial guidance system (ICS) was assumed 

to be as reliable as the booster IGS; 

9) 2 CS-156 SRM reliability data would be derived from 

the 3 CS-156 SRM reliability data. 

B. APPROACH 

To achieve the study objectives, the analyses were conducted 

within the failure mode and effects program, which consists of 

four basic steps: 

1) Identification of blackbox failure modes; 

2) Evaluation of end effects on the booster as a result 

of these failure modes; 

3) Assessment of the probability of occurrence of these 

failure modes; 

4) Data reduction' and analyses. 

The six performance configurations defined in Chapter II and 

the Titan IIIC configuration are subdivided into basic systems by 

associate contractors and are discussed in this section. 

1. Aeroiet-General Corporation 

The analyses provided by Aerojet included failure mode identification and the probabilities of occurrence, which were based on 

measured data from test firings. Inputs covered: 

1) Stage I (8:1) - A Titan IIIC Stage I engine without 

modification; 

2) Stage I (15:1) - An updated engine including performance and reliability modifications; 

3) Stage II - A Titan IIIC Stage II engine without modification; 
NRO APPROVED FOR 

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IX-3 

4) Transtage - A Titan IIIC transtage engine without 

modification. 

The analyses of end effects on the booster were completed by the 

Martin Company. 

2. AC Electronic Division (ACED) 

The analyses provided by ACED included failure mode identification, end effect analyses, and probabilities of occurrence, 

which were based on measured failure rates and updated environmental factors. Inputs included a booster IGS for with-transtage 

configurations, which included pressure temperature control (PTC), 

and for without-transtage configurations, which did not include a 

PTC. 

d Tech 

The analyses provided by UTC included failure mode identification and the probabilities of occurrence, which were based 

on failure rate data from failure rate handbooks, degradation 

or use factors, and updated environmental factors. Inputs included data on: 

1) 5 seg 120 SRM - The Titan IIIC SRM and TVC.system 

without modification; 

2) 7 seg 120 SRM - The modified Titan IIIC SRM and a 

baseline redundant TVC system; 

3) 3 CS-156 SRM - A new SRM and a baseline redundant 

TVC system. 

Martin Company 

The analyses conducted by Martin Company were based on two 

hardware configurations, i.e., the with-transtage and the withouttranitage core configurations. Failure mode identification and 

vehicle effect analyses were completed for each blackbox within 

these configurations. To assess the probability of occurrence 

for each failure mode, failure rates were obtained from handbook 

and test data and subsequently degraded to achieved failure rates 

using Titan IIIC background data. Updated environmental factors 

for each of the six performance configurations (based on recent 

Titan IIIC flight teat data) and mission time profiles were programed into an IBM 1620 computor for computation of probability 
NRO APPROVED FOR 

RELEASE 1 JULY 2015 

IX-4 

of occurrences for all six performance configurations in accordance with the formula 

P

f =nxtxK

op 

 x f

r 

where, 

n = quantity of blackboxes, 

t = time of operation (hr), 

Kop = environmental factor, 

f

r 

= failure rate in expected number of occurrences/ 

million hours. 

All the associate contractor and Martin inputs were prepared 

on electronic data-processing input transmittals, which were subsequently key-punched on IBM cards for mechanized reduction. The 

IBM 1620 program mentioned above was used to obtain assessments 

of mission success and crew safety for each subsystem for each 

of the six performance configurations. The Titan IIIC configuration mission success assessment was determined from the withtranstage core configurations by eliminating the redundancy and 

switchover modifications peculiar to the performance modifications. 

C. RESULTS 

The results of the reliability and crew safety study are summarized in Tables IX-1 thru IX-4. 

Crew Safety,/Mission Abort Summary 

Table IX-1 is a configuration comparison summary by warning 

time, total number of aborts, and mode IV aborts. Warning time 

analysis indicates that the best crew safety configuration is the 

7 seg-120 configuration (without transtage), although there is 

only a difference of 107 in the number of mission aborts between 

the without-transtage configurations. The 7 seg-120 configuration 

has the least mission aborts with less than 3 sec warning time 

for the with-transtage configurations, although there is only a 

difference of 147 between the with-transtage configurations. Each 

SRM configuration without the transtage has approximately 10% 
NRO APPROVED FOR 

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IX-5 

fewer mission aborts with less than 3 sec jead time than the corresponding SRM configuration with the transtage. The expected 

number of mode IV aborts are the number of aborts that would result in Gemini splashdown below the 30th parallel. 

Table IX-1 Crew Safety/Mission Aborts Summary 

With 

7 seg120 

Transtage 

2 CS156 

3 CS156 

7 seg120 

Without 

2 CS15'6 

Transtage 

3 CS156 Titan IIIC 

Expected Number 

(F) of Aborts 

by Warning Time 

0 < F

1 

< 3 sec 2,401 2,451 2,548 2,202 2,222 2,309 

3 sec < F2 <6 

sec (40 to 90 sec) 66 65 66 66 65 65 

6 sec < F3 35,978 37,976 38,420 31,782 33,788 34,138 -- 

Not Analyzed 2,042 2,953 3,078 2,041 2,953 3,078 

Total Number of 

Aborts 40,484 43,442 44,008 36,099 39,025 39,586 56,080 

Expected Number 

of Mode IV Aborts 570 681 773 -- 

Note: Figures represent occurrences in PPM. 

2. Mission Aborts by Flight Phase 

Table IX-2 summarizes mission aborts by flight phase for each 

configuration where flight phases are defined as follows: 

1) Stage 0 flight phase - The interval from SRM ignition 

to Stage I Start; 

2) Stage I flight phase - The interval from Stage I start 

.to Stage I/Stage II staging; 
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RELEASE 1 JULY 2015 

3) Stage II flight phase - The interval from Stage I/ 

Stage II staging to Stage III start or payload separation in the case of the without-transtage configurations; 

4) Transtage burn 1 - The interval from transtage start 

thru shutdown; 

5) Transtage coast - A 45-minute coast period; 

6) Transtage burn 2 - The period from transtage start 

thru payload separation. 

Analysis of Table IX-2 indicates that the 7 seg-120 configuration without the transtage has the least number of mission aborts 

for the 80-n-mi orbit. The 7 seg-120 configuration with the 

transtage shows only a difference of 2464 in mission aborts between the 80- and 130-n-mi orbits. 

Table IX-2 Mission Aborts by Flight Phase 

FLIGHT PHASE 

With Transtage Without Transtage 

Titan IIIC 

7 seg120 

2 CS156 

3 CS156 7 seg120 

2 CS156. 

3 CS156 

Stage 0 8,058 9,404 10,124 

15,101 

8,058 

13,133 

9,404 10,124 10,495 

Stage I 13,642 15,265 14,709 

14,913 

- 

-- 

-1--- 

14,550 17,000 

Stage II 14,912 14,915 

1,409 

40,993 

14,905 

1,409 

14,908 

-- 

14,912 16,671 

3,983 

Transtage (1st 

burn) 

Subtotal 

(80-n-mi equivalent) 

1,409 

38,022 41,539 36,099 39,025 39,586 48,149 

._. 

Transtage (coast) 

,_ 

478 463 488 

_____ _ 

5,867 

2,064 

Transtage (2nd 

burn) 1,986 1,986 1,986 

Subtotal 2,464 2,449 2,469 7,931 

Total 

(130-n-mi equivalent) 40,484 43,442 44,008 56,080 
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3. Crew Safety/Mission Abort Summary by Subsystem Grouping, 

Table IX-3 is a configuration comparison by subsystem grouping 

defined as follows: 

1) Total system; 

2) Propulsion, which includes LREs and the SRMs less the 

TVC; 

3) Guidance and controls, which include the guidance, 

flight controls, and the TVC subsystems; 

4) Other, which includes the remainder of the subsystems 

not listed in 2) and 3). 

The 7 seg-120 configuration is considered baseline and the other 

configurations are shown as deltas from this baseline. 

Table IX-3 Crew Safety/Mission Abort Summary by Subsystem 

Grouping 

With Transtage Without Transtage 

7 seg120 

2 CS1566 

3 CS156A 7 seg120 

2 CS156A 

3 CS1564 

Total

System 

R 40,484 2,958 3,524 -4,385 -1,459 -898 

F

1 2,401 50 147 -199 -179 -92 

F

2 66 -1 0 0 -1 -1 

Propulsion 

R 34,431 2,902 3,446 -2,702 209 745 

F1 1,372 57 151 -91 -34 59 

F2 0 0 0 0 0 0 

Guidance 

and 

Control 

R 

F

1 

1,402 53 77 -4 52 78 

97 13 13 7 22 26 

F

2 66 -1 -1 0 -1 -1 

Other 

R 4,651 10 1 -1,679 -1,720 -1,721 

F

, 1 932 -20 17 -115 -167 -177 

2 0 0 0 0 0 
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1X-8 

Effect of Step 0 Burning Time on Reliability 

Table 1X-4 is intended to show the increase or decrease in 

mission aborts as a result of increased burning time during Step 

0. The 7 seg-120 configuration is considered baseline and the 

other configurations are shown as deltas from the baseline. Step 

0 is defined as the mission flight time from Stage 0 ignition 

through SRM staging. The core is defined as all subsystems excluding LREs, which are not sensitive to the SRM burning time. 

Table IX-4 Effects of Step 0 Burning Time on Reliability 

------------ 

Without Transtage 

Configuration 7 seg-120 2 CS-156 3 CS-156 7 seg-120L 2 CS-15 3 CS-156L\ 

Burn Time 120 sec 136 sec 150 sec 120 sec 136 sec 150 sec 

SRM 7328 1407 1942 0 1407 1943 

-23 

75 

Core 1659 219 90 

75 

-113 

0 

106 

TVC 811 53 53 

Total 

Aborts 

9798 1679 2107 -113 1566 1994 

With Transtage 
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X-1 

X. SCHEDULE 

The schedule and program plans provide a basis for evaluating 

relative schedule position of the six configurations under study 

(see Table II-1 and Fig. II-1 for a description of the configurations). 

To provide the schedule comparison, a Titan III/MOL booster 

program plan was established. The six configurations, based on 

an equal-risk schedule, are then compared against the program 

plan. 

A. GROUND RULES 

Planning for the booster and Stage 0 development programs is 

based on the following ground rules and restraints: 

1) The core booster subsystems schedules are based on 

the reference configuration described in Chapter II; 

2) Phase II go-ahead for the LRE, SRM, and guidance 

associate contractors will be the same as the core 

booster; 

3) The Integrated Launch Complex (ILC) will be used for 

MOL launches. This facility will be activated before 

the MOL schedule requirements and will not constrain 

launch capability; 

4) Resource application (facilities, manpower, number 

of shifts, overtime, etc) will be on the same basis 

as the Titan III development program; 

5) Booster acceptance tests will be performed in Cell 

P-4, This cell will be modified and equipped to conduct test on the Titan IIIC configuration and the selected MOL configuration. 
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B. TITAN III/MOL BOOSTER PROGRAM PLAN 

The program plan, Fig. X-1, was developed to establish a baseline to compare schedule and cost of the various configurations. 

This plan supports a launch schedule established by the Space Systems Division of the Air Force Systems Command. 

The significant elements are: 

1) Program go-ahead on 1 March 1965; 

2) Airborne engineering and major test comprised of 

structural, component development, controls mockup, 

and design assurance; 

3) Airborne first article fabrication, which covers the 

modification of Vehicle 17 from the Titan IIIC configuration, and the first production MOL; 

4) Acceptance test of the seven MOL configuration vehicles will be conducted in Cell P-4, Vehicle 17 will 

be the first MOL booster accepted in November 1967 

(the additional vehicles to be tested in Cell. P-4, 

as reflected in Fig. X-1, are Titan IIIC follow-on 

vehicles displayed for a cost base only per SSD direction); 

5) The CFP delivery requirements for. LREs, SRNs, and 

ICS to support the MOL booster fabrication and launch 

program are indicated on the plan; 

6) The ILC design fabrication, nstallation, and acceptance plan is outlined on the program plan. This plan 

calls for a Phase II go-ahead of 1 February 1966. 

The activation and acceptance of the launch facility 

cheduled in October 1967; 

7) There will be seven MOL launches from ILC. The first 

launch is planned for May 1968, and launches will 

continue at a rate of one every four months through 

May 1970. 
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UMMMWWWWWWIEW WNW= 1W. .WIIMUU 

=MOM WrWmilimulW OM OWM WINWERWEIMI lmm 

itan IIItMOL Booster Program Plan 

011111 MI MO MEM WM Ma= W 

WHO ti 

M MO. OmOOMMOWIAMEMOLIMOMMOOMMIMEM 

SSD-CR-65-206 
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X-4 

C. EQUAL RISK SCHEDULE 

A schedule review of the six configurations was made on a 

basis of equal risk from program go-ahead through the first launch. 

To establish a schedule of equal risk, a standard plan for 

each configuration was established. The plan was then converted 

into a PERT network where each element was evaluated. The evaluation considered such items as, new design, production capability, 

number of tests required, probability of success, etc. Each evaluation was made on an optimistic, most likely, and pessimistic 

basis to establish span time for that effort. Span time on all 

the networks is expressed in weeks from go-ahead. We concluded 

that this method was the most adequate way to demonstrate equal 

risk comparison between various configurations, 

The elements of the standard network are: 

1) Airborne core subsystem design; 

2) Engineering development and design assurance test; 

3) Structural test; 

4) Fabrication and delivery of first article; 

5) Launch facility construction; 

6) Airborne ground equipment design, fabrication, installation, and activation; 

7) LRE development and delivery; 

8) SRM development and delivery; 

9) Launch of first MOL. 

:-12O with Trans a e. 15:1 Stare I ation 

The schedule for this configuration is reflected in Fig. X-2. 

The expected span time from go-ahead to launch for this configuration is 117.1 weeks, Based on a Phase II go-ahead of 1 March 

1966, the launch would be 19 June 1968. 
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T1CAL PATH 

FE Ci

2 

U N 

co

l 

7 seg-120 w ith Transtage; 15:1 Stage I Engine Con figruation Sc 

S S D -CR - 61) - 20 6 
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X-6 

The most critical path is the LRE development test, production, LRE on dock at Martin, and completion of the first MOL 

booster installation. This path controls the booster acceptance 

test and subsequent launch. 

The second most critical path is through the guidance and 

control new design to first MOL'booster installations complete. 

This is caused by the new design and development of actuators. 

An alternative plan, although of some higher risk, could be 

followed that has been used in the past. This plan is shown in 

Fig. X-2 and calls for the use of a checkout engine for booster 

acceptance test and installation of the flight engine before shipment to the launch site. The alternative plan moves the launch 

up to 3 April 1968. 

2. 7 se 120 without Transta ' 15: Sta e En ne Con i ura ion 

The schedule of all configurations showed no significant 

change when the transtage was removed. Note in Fig. X-3 that 

this configuration has the same launch capability as the previous 

configuration and the same critical paths. 

There is some reduction in the structural test span time as a 

result of fewer tests. The requirement to develop new transtage 

actuators would also be e,liminated. 

The same alternative plan exists for this configuration as 

described for the configuration discussed in Subsection 1 above. 

3,2ssLL1§. with Transta ' 8:1 Sta e En ne Conf 

The schedule for this configuration is reflected in Fig. X-4. 

The expected span time from go-ahead to launch is 121 weeks. 

Based on a I March 1966 Phase II go-ahead, the launch would be 

19 July 1968. 

The most critical path is the SRM development firings and SRN 

preflight rating test complete. This path controls the first 

launch because of the requirement to complete all PFRTs before 

launch. 

The second most critical path is the flight ORM on dock at 

the launch site. This controls the launch pad checkout and launch. 

A more detail analysis of the SRN development schedule is presented in Section D. The critical elements of the SRM development 

schedule are also covered in Section D. 
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X-8 

LEGEND: SPAN TIME IN WEEKS 

a. 

6 

2 CS-156 with Transtage; 8:1 StageI EngineConfigura tion Schedu le 
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X-9 

2 CS-156 without Transta 8:1 Stage I En ine Con i uration 

The schedule for this configuration is the same as the preceding one and is shown in Fig. X-4. 

5. 3 CS-156 with TranstageL 8:1 Stage I Engine Configuration 

The schedule for this configuration is shown in Fig. X-5. 

The expected span time from go-ahead to launch is 123 weeks. 

Based on a 1 March 1966 Phase II go-ahead, the launch would be 

19 July 1968. 

The most critical path is the same as the 2 segment. This 

path controls the first launch because of the requirement to complete all PFRTs before launch. 

The second most critical path is the flight SRM on dock at the 

launch site. This controls the launch pad checkout and launch. 

A more detailed analysis of the SRM development schedule is presented in Section D. The critical elements of the SRM development schedule are also covered in Section D. 

6 3 CS-156 without Transta e. 8:1 Stage I Engine Configuration 

The schedule for this configuration is the same as the preceding one and is shown in Fig. X-5, 
3 CS-156 with Transtage; 8:1 Stage I Engine Configuration Schedule 

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D. EQUAL RISK SCHEDULE FOR STAGE 0 DEVELOPMENT 

The Stage 0 development is a significant factor in the total 

MOL booster program. To obtain a better evaluation of the impact 

on the first launch capability, an equal-risk schedule was made 

for the three SRM configurations. 

The schedule spans represent a composite of data obtained 

from United Technology Center, Lockheed Propulsion Company, Thiokol 

Chemical Corporation, Aerojet-General Corporation, and data from 

the 5-segment Titan IIIC program. 

The basic elements of the standard Stage 0 development plan 

are: 

1) Case design and procurement; 

2) TVC design and procurement; 

3) Nozzle design and procurement; 

4) Subscale test; 

5) Thrust termination test; 

6) AGE design, fabrication, installation, and checkout; 

7) Development firings; 

8) PFRT firings; 

9) Assembly and checkout of SRMS at launch site, 

1, 7 seg-120 SRM Develomnt Schedule 

The schedule for this configuration is reflected in Fig. X-6. 

The'span time from go-ahead to launch is 94.2 weeks. Based on a 

1 March 1966 go-ahead, the launch would be 10 January 1968. It 

is significant to note that the launch capability is much earlier 

for this configuration than that for the 7 seg-120 with- and without-transtage configurations. In the previous configurations, 

the critical path was through booster availability as a result of 

liquid rocket engine development, where this configuration launch 

capability is not restrained by the core booster. 
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X-12 SSD-CR-65-206 

LEGEND: SPAN TIME IN WEEKS 

CRITICAL PATH 

U 

a 

7 seg-120 SRM Development Schedule 
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X-13 

The development span time is shorter for the 7 seg-120 SRMs 

than for the 2 and 3 CS-156 SRMs, Three items primarily contribute to the shorter spans: 

1) Less new design, 

2) Fewer development firings, 

3) Less procurement time for the motor cases. 

The most critical path for this configuration is through 1st 

case segment available, 1st development motor available, development test, and PFRT test. This path controls the first launch 

capability; however, this is considerably ahead of the program 

plan launch requirement, 

The second most critical path is the nozzle design and availability of the 1st development motor. 

2 CS-156 SRM Develo Schedule 

The schedule for this configuration is shown in Fig. X-7. The 

span time from go-ahead to launch is 121 weeks. Based on a 1 March 

1966 go-ahead, the launch would be 19 July 1968, 

The most critical path is from the 1st motor case available, 

1st development motor available, development firings, and PFRT 

tests complete. This path controls 1st launch as PFRT tests must 

be completed prior to launch. 

The second most critical path is from nozzle design through 

availability for the 1st development motor fabrication, 

The launch capability for this configuration, based on equal 

risk, does not meet the May 1968 launch schedule of the program 

plan, To meet the launch schedule, the following program alternatives are available; 

1) Long lead go-ahead for case material and nozzle design 

could reduce the span time sufficiently from Phase II 

go-ahead to meet the launch requirement. This is one 

of the most desirable approaches as it does not increase risk; 
2 CS-156 SRM Development Schedule 

SPAN TIME IN WEEKS 

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2) Reduce PFRT requirements for the first launch. This 

could.be obtained by conducting fewer firings or by 

reducing the number of successful PFRT firings required for launch. A number of successful firings on 

production motors could still be demonstrated prior 

to launch, but this plan would be of some higher risk; 

3) Decrease the number of development firings prior to 

PFRT firings. This action would reduce span time to 

meet launch, but would be of greater risk because of 

the reduced development firings before production 

motor test. 

3 CS-156 SRM Development Schedule 

The schedule for this configuration is shown in Fig. X-8. 

The span time from go-ahead to launch is 123 weeks. Based on a 

1 March 1966 go-ahead for Phase II, the launch capability would 

be 2 August 1968. The only difference between the 2 and 3 CS-156 

SRM configurations is the estimated two weeks span to fabricate 

an additional segment for the first development motor. The same 

critical paths exist for both configurations. The schedule difference, on an equal risk basis, is insignificant between the two 

configurations. 

To meet the May 1968 launch schedule, the program alternatives would be the same.' 
x-16 

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SPAN TIME N I WEEKS 

0 

CRITICAL PATH 

—— 2ND MOST CRITICAL PATH 

Development Schedule 

SSjiCR6: 206 
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X-17 

E. SUMMARY 

From a schedule reference, as a result of the performance 

improvement study, the following conclusions can be drawn: 

1) The 7 seg-120, 15:1 Stage I engine configuration is 

the least schedule risk program; 

2) The 2 and 3 CS-156, 8:1 Stage I engine configurations 

are equal schedule risk programs; 

3) Configurations with or without transtage are not significantly different in schedule risk; 

4) All configurations with the appropriate long lead 

go.ahead can meet the May 1968 launch schedule with 

equal risk. 
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SSD-CR-65-206 XI-1 

XI, BIBLIOGRAPHY 

Martin Company Interface Control Drawings 

IGS Physical Interface, ICD-IIII-11001 

IGS Installation Interface, ICD-TIII-11003 

Stage I Engine Physical Interface, ICD-TIII-12101 

Stage II Engine Physical Interface, ICD-TIII-12201 

SRM Physical Interface, ICD-TIII-14001 

Martin Company Titan III/MOL Compatibility Study Reports 

Evaluation of Design Characteristics of 7-Segment, 156-Inch 

2-Segment, and 156-Inch 3-Segment SRMs, K3-002, 10 September 

1965 

Stability Effects of Each Performance Improvement Configuration, K3-031 (to be issued) 

Airborne Structures Configuration, K3-045, 23 August 1965 

Analysis of Structural Loads, K3-501, 15 September 1965 

Stage I Longeron Reanalysis, K3-503, 10 September 1965 

Improved Engine, Titan III/MOL Stage I, K3-514, 20 August 

1965 

Systems Design Criteria, K3-557, 4 August 1965 

Panel Flutter Study, K3-561, 24 August 1965 

Analysis in Support of Core Modifications, K3-570, 23 August 

1956 

Miscellaneous Documents 

156 SRM Strap-On Stage Zero for Improved Titan III Performance, 

LPC TR P-2467, 15 August 1965, Lockheed Propulsion Company 

3-Segment 156-Inch Diameter Motor Final Design and Performance 

Summary, UTC 4801-65-3, 6 August 1965, United Technology 

Center 
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XI-2 SSD-CR-65-206 

156-Inch SRM MOL Study, Appendix I & II, ICC DLB-085-PD, 2 

August 1965, Thiokol Chemical Corporation 

Technical Data on u ort o MOL/Titan III Booster 156-Inch 

S Stage O. AGC SNP-65714, 13 August 1965, Aerojet-General 

Corporation 

"Titan III/MOL Data Book." TM 5141/31-65-16 (Rev 1), Martin 

Company, Denver, Colorado, 27 August 1965. 

Detail Specification for Standard Space Launch Vehicle, Standard SSS-TIII-D-SLV/01A010, Martin Company, Denver, 

Colorado, 

HuMan Engineering Design Criteria. MIL-STD-803A, Department 

of the Air Force, 27 January 1964. 

"Titan III (SSLV 624A) Solid Rocket Motor Staging Analysis," 

TM 5141/33-64-17, Martin Company, Denver, Colorado, May 

1964. 

"Stage I NPSH Requirements for MOL." (Letter from D. E. 

Robison, AGC, to L. J. Adams, Martin Company), 24 August 

1965. 

"Propulsion Performance Data for MOL Study." (TWX .rom L. 

Beard, Aerospace to L. J. Adams, Martin Company), 27 July 

1965, 

"Titan III/MOL Stage I and II Class Nominal Engine Performance." 

(IDC from Martin Company Powerplants Unit to L. J. Paustian 

20 July 1965. 
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SSD-CR-65-206 

DISTRIBUTION 

21 thru 30 

31 thru 80 

Headquarters 

Space Systems Division 

Air Force Systems Command 

Air Force Unit Post Office 

Los Angeles, California 90045 

Attn: SSBE 

Defense Documentation Center (DDC) 

Cameron Station 

Alexandria, Virginia 

Martin Company 

Denver Division 

Denver, Colorado 80201 

Attn: Classified Loan Center 

C. o p 

thru 20 
VEZETTgo7 

ra 
1 

'42`.17=7,1 
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14. 

; 

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7_:i 

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4 

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