Saturn 5 Launch Vehicle Flight Evaluation Report: AS-506, Apollo

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GEORSE 8.,

MPR- SAT- FE- 69-9 SEPTEMBER 20, 1969

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SATURN V LAUNCH VEHICLE

FLIGHT EVALUATION REPORT-AS-506

APOLLO !1 MISSION _3 ('NASA-_ N-X-625:58) SATURN L _I U t_'_C !I_ VkHICLE _'4'? O- 70 _ :"3i F t. I O•H T _: V A L U A T I_0 i_I_ _'__'_p i3 i::t.T : A S- 5 0 5 A.'v'i" ___L ,.I_"-_ i 1 MIS.,.S[L._ (NASA) Z6# p u n c I a s _O/Z _' "_ _ ___'_ 74

PREPARED BY "'_,,'-// - C "%"_ SATURN FLIGHT EVALUATION WORKING GR_h _ )_.'}:"_

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

_C - Form 774 (Rev October 1967) GEORGE C. MARSHALL SPACE FLIGHT CENTER

MPR-SAT-FE-69- 9

SATURNV LAUNCH VEHI CLE FLIGHT EVALUATION REPORT-AS-.506

APOLLO 11 MI SSION

PREPARED BY SATURN FLIGHT EVALUATION WORKING GROUP _i_ _:!:!iiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiii_iii!iiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiii!!_!_::i_!ii_:!_i:_>:_

AS-506 LAUNCH VEHICLE MPR-SAT-FE-69-9

SATURN V LAUNCH VEHICLE FLIGHT EVALUATION REPORT - AS-506

APOLLO II MISSION

Saturn Flight Evaluation Working Group George C. Marshall Space Flight Center

ABSTRACT

Saturn V AS-506 (Apollo II Mission) was launched at 09:32:00 Eastern Daylight Time on July 16, 1969, from Kennedy Space Center, Complex 39, Pad A. The vehicle lifted off on schedule on a launch azimuth of 90 degrees east of north and rolled to a flight azimuth of 72.058 degrees east of north.

The launch vehicle successfully placed the manned spacecraft in the planned translunar injection coast mode. The S-IVB/IU was placed in a solar orbit with a period of 342 days by a combination of continuous LH2 vent, a LOX dump and APS ullage burn.

The Principal and Secondary Detailed Objectives of this mission were completely accomplished. No failures, anomalies, or deviations occurred that seriously affected the flight or mission.

Any questions or comments pertaining to the information contained in this report are invited and should be directed to:

Director, George C. Marshall Space Flight Center Huntsville, Alabama 35812 Attention: Chairman, Saturn Flight Evaluation Working Group, S&E-CSE-LE (Phone 453-2575) TABLEOF CONTENTS

Section Page

TABLEOF CONTENTS iii LIST OF ILLUSTRATIONS xi LIST OF TABLES xvi ACKNOWLEDGEMENT xix

ABBREVIATIONS XX MISSION PLAN xxiii

FLIGHT TESTSUMMARY XXV

INTRODUCTION I.I Purpose 1.2 Scope EVENTTIMES 2.1 Summaryof Events 2-I 2.2 Variable Time and CommandedSwitch Selector Events 2-2

LAUNCHOPERATIONS 3.1 Summary 3-I 3.2 Prel aunch Milestones 3-I 3.3 Countdown Events 3-I 3.4 Propella'nt Loading 3-I 3.4.1 RP-I Loading 3-I 3.4.2 LOX Loading 3-3 3.4.3 LH2 Loading 3-3 3.4.4 Auxiliary Propulsion System Propellant Loading 3-4 3.5 S-II Insulation, Purge and Leak Detection 3-4

iii TABLEOF CONTENTS(CONTINUED)

Section Page 3.6 Ground Support Equipment 3-4 3.6.1 Ground/Vehicle Interface 3-4 3.6.2 MSFCFurnished Ground Support Equipment 3-5 3.6.3 Camera Coverage 3-6

4 TRAJECTORY

4.1 Summary 4.2 Tracking Data Utilization 4.2.1 Tracking During the Ascent Phase of F1 i ght 4-2 Tracking During Orbital Flight 4-2 Tracking During the Injection Phase of F1 ight 4-2 4.2.4 Tracking During the Post Injection Phase of Flight 4-3

4.3 Trajectory Evaluation 4-3 4.3.1 Ascent Trajectory 4-3 4.3.2 Parking Orbit Trajectory 4-6 4.3.3 Injection Trajectory 4-6 4.3.4 Post TLI Trajectory 4-10 4.3.5 S-IVB/IU Post Separation Trajectory 4-11

S-IC PROPULSION

5.1 Summa ry 5-I

5.2 S-IC Ignition Transient Performance 5-I 5.3 S-IC Mainstage Performance 5-3

5.4 S-IC Engine Shutdown Transient Performance 5-5

5.5 S-IC Stage Propellant Management 5-6 5.6 S-lC Pressurization Systems 5-7 5.6.1 S-IC Fuel Pressurization System 5-7 5.6.2 S-IC LOX Pressurization System 5-7

5.7 S-IC Pneumatic Control Pressure System 5-8

5.8 S-IC Purge Systems 5-8

5.9 S-IC POGO Suppression System 5-9

iv TABLE OF CONTENTS (CONTINUED)

Section Page

6 S-II PROPULSION

6.1 Summary 6-I 6.2 S-II Chilldown and Buildup Transient Performance 6-2

6.3 S-II Mainstage Performance 6-4

6.4 S-II Shutdown Transient Performance 6-8

6.5 S-II Stage Propellant Management 6-8 6.6 S-II Pressurization Systems 6-9 6.6.1 S-II Fuel Pressurization System 6-9 6.6,2 S-II LOX Pressurization System 6-11

6.7 S-II Pneumatic Control Pressure Sys tern 6-13 6.8 S-II Helium Injection System 6-13

7 S-IVB PROPULSION

7.1 Summary 7-I 7.2 S-IVB Chilldown and Buildup Transient Performance for First Burn 7-I

7.3 S-IVB Mainstage Performance for First Burn 7-3

7.4 S-IVB Shutdown Transient Performance for First Burn 7-5

7.5 S-IVB Parking Orbit Coast Phase Condi ti oni ng 7-5 7.6 S-IVB Chilldown and Restart for Second Burn 7-8

7.7 S-IVB Mainstage Performance for Second Burn 7-12

7.8 S-IVB Shutdown Transient Performance for Second Burn 7-12

7.9 S-IVB Stage Propellant Management 7-12

7.10 S-IVB Pressurization System 7-14 7.10.1 S-IVB Fuel Pressurization System 7-14 7.10.2 S-IVB LOX Pressurization System 7-19 TABLEOF CONTENTS(CONTINUED)

Section Page 7.11 S-IVB Pneumatic Control System 7-25 7.12 S-IVB Auxiliary Propulsion System 7-25 7.13 S-IVB Orbital Safing Operations 7-27 7.13.1 Fuel Tank Safing 7-27 7.13.2 LOXTank Dumpand Safing 7-27 7.13.3 Cold Helium Dump 7-29 7.13.4 Ambient Helium Dump 7-29 7.13.5 Stage Pneumatic Control Sphere Safing 7-29 7.13.6 Engine Start Sphere Safing 7-31 7.13.7 Engine Control Sphere Safing 7-31 8 HYDRAULICSYSTEMS 8.1 Summary 8-I 8.2 S-IC Hydraulic System 8-I 8.3 S-II Hydraulic System 8-I 8.4 S-IVB Hydraulic System 8-2 STRUCTURES 9.1 Summary 9-I 9.2 Total Vehicle Structures Evaluation 9-I 9.2.1 Longitudinal Loads 9-I 9.2.2 Bending Moments 9-2 9.2.3 Vehicle Dynamic Characteristics 9-3 9.2.3.1 Longitudinal Dynamic Characteristics 9-3 9,2.3.2 Lateral Dynamic Characteristics 9-12 I0 GUIDANCEANDNAVIGATION 10. 1 Summary I0-I I0.I,I Flight Program I0-I 10.1.2 Instrument Unit Components I0-I 10.2 Guidance Comparisons I0-I 10.2.1 Late S-II Stage EMRShift 10-4 10.3 Navigation and Guidance Scheme Evaluation 10-9 10.4 Guidance System Component Evaluation I0-I0 10.4.1 LVDCPerformance I0-I0 10.4.2 LVDA Performance I0-I0 10.4.3 Ladder Outputs I0-I0

vi TABLE OF CONTENTS(CONTINUED)

Secti on Page I0.4.4 Telemetry Outputs I0-I0 10.4.5 Discrete Outputs I0-13 I0.4.6 Switch Selector Functions I0-13 10.4.7 ST-124M-3 Inertial Platform I0-14

II CONTROLSYSTEM II .I Summary II-I II .2 S-IC Control System Evaluation II-I II ,2.1 Liftoff Clearances 11-2 II ,2.2 S-lC Flight Dynamics 11-2 II .3 S-II Control System Evaluation 11-6 II .4 S-IVB Control System Evaluation 11-13 II .4.1 Control System Eval uation During First Burn 11-13 II ,4.2 Control System Evaluation During Parking Orbit 11-13 II .4.3 Control System Evaluation During Second Burn 11-14 II .4.4 Control System Evaluation After S-IVB Second Burn 11-15

12 SEPARATION 12.1 Summary 12-I 12.2 S-IC/S-II Separation Evaluation 12-I 12.3 S-II/S-IVB Separation Evaluation 12-I 12.4 S-IVB/IU/LM/CSM Separation Evaluation 12-I 12.5 Lunar Module Docking and Ejection Evaluation 12-2

13 ELECTRICALNETWORKS 13.1 Summary 13-I 13.2 S-IC Stage Electrical System 13-I 13.3 S-II Stage Electrical System 13-2 13.4 S-IVB Stage Electrical System 13-3 13.5 Instrument Unit Electrical System 13-6

14 RANGESAFETYAND COMMANDSYSTEMS 14.1 Summary 14-I 14.2 Secure Range Safety CommandSystems 14-I

vii TABLEOF CONTENTS(CONTINUED)

Section Page 14.3 Commandand Communications System 14-I 15 EMERGENCYDETECTIONSYSTEM 15.1 Summary 15-I 15.2 System Evaluation 15-I 15.2.1 General Performance 15-I 15.2.2 Propulsion System Sensors 15-I 15.2.3 Flight Dynamics and Control Sensors 15-I 16 VEHICLEPRESSUREENVIRONMENT 16.1 Summary 16-I 16.2 Base Pressures 16-I 16.2,1 S-IC Base Pressures 16-I 16.2.2 S-II Base Pressures 16-I 16.3 Surface Pressure and Compartment Venting 16-5 16.3.1 S-IC Stage 16-5 16.3.2 S-II Stage 16-5 17 VEHICLETHERMALENVIRONMENT 17.1 Summary 17-I 17.2 S-IC Base Heating 17-I 17.3 S-II Base Region Environment 17-4 17.4 Vehicle Aeroheating Thermal Environment 17-7 17.4.1 S-IC Stage Aeroheating Environment 17-7 17.4.2 S-II Stage Aeroheating Environment 17-8 18 ENVIRONMENTALCONTROLSYSTEM 18.1 Summary 18-I 18.2 S-IC Environmental Control 18-I 18.3 S-II Environmental Control 18-2 18.4 IU Environmental Control 18-2 18.4.1 Thermal Conditioning System 18-2 18.4,2 Gas Bearing Supply System 18-7 19 DATASYSTEMS 19.1 Summary 19-I 19.2 Vehicle Measurement Evaluation 19-I 19.3 Airborne Tel emetry Systems 19-2

viii TABLEOF CONTENTS(CONTINUED)

Section Page 19.4 RF Systems Evaluation 19-6 19.4.1 Telemetry System RF Propagation Evaluation 19-6 19.4.2 Tracking Systems RF Propagation Evaluation 19-6 19.4.3 CommandSystems RF Evaluation 19-8 19.5 Optical Instrumentation 19-12 2O MASSCHARACTERISTICS 20.1 Summary 20-I 20.2 Mass Evaluation 20-I

21 MISSION OBJECTIVESACCOMPLISHMENT 21-I

22 FAILURES, ANOMALIESANDDEVIATIONS 22.1 Summary 22-I 22.2 System Failures and Anomalies 22-I 22.3 System Devi ati ons 22-I 23 SPACECRAFTSUMMARY 23-I

Appendix A ATMOSPHERE A.I Summary A-I A.2 General Atmospheric Conditions at Launch Time A-I A.3 Surface Observations at Launch Time A-I A.4 Upper Air Measurements A-I A.4.1 Wind Speed A-I A.4.2 Wind Direction A-I A.4.3 Pitch Wind Component A-2 A.4.4 Yaw Wind Component A-2 A.4.5 Component Wind Shears A-2 A.4.6 Extreme Wind Data in the High Dynamic Region A-3 A.5 Thermodynamic Data A-3 A.5.1 Temperature A-3 A.5.2 Atmospheric Pressure A-IO

ix TABLE OF CONTENTS (CONTINUED)

Appendix Page

A.5.3 Atmospheric Density A-IO A.5.4 Optical Index of Refraction A-13

A.6 Comparison of Selected Atmospheric Data for Saturn V Launches A-13

B • AS-506 SIGNIFICANT CONFIGURATION CHANGES

B. 1 I ntroducti on B-I LIST OF ILLUSTRATIONS

Figure Page 2-I Telemetry Time Delay 2-2

4-I Ascent Trajectory Position Comparison 4-3 4-2 Ascent Trajectory Space-Fixed Velocity and Flight Path Angle Comparisons 4-4 4-3 Ascent Trajectory Acceleration Comparison 4-5

4-4 Dynamic Pressure and Mach Number Comparisons 4-6

4-5 Ground Track 4-12 4-6 Injection Phase Space-Fixed Velocity and Flight Path Angle Comparisons 4-13 4-7 Injection Phase Acceleration Comparison 4-14

4-8 Slingshot Maneuver Longitudinal Velocity Increase 4-15 4-9 Trajectory Conditions Resulting from Slingshot Maneuver Velocity Increments 4,.18 4-10 S-IVB/IU Velocity Relative to Earth Distance 4-18

5-I S-IC LOX Start Box Requirements 5-2 5-2 S-IC Engines Buildup Transients 5-3

5-3 S-IC Stage Propulsion Performance Parameters 5-4 5-4 S-IC Fuel Ullage Pressure 5-8

5-5 S-IC LOX Tank Ullage Pressure 5-9 5-6 S-IC LOX Suction Duct Pressure, Engine No. 5 5-10

6-I S-II Engine Start Tank Performance 6-3

6-2 S-II Engine Pump Inlet Start Requirements 6-5 6-3 S-II Steady-State Operation: 6-6

6-4 S-II Fuel Tank Ullage Pressure 6-10 6-5 S-II Fuel Pump Inlet Conditions 6-12

6-6 S-II LOX Tank Ullage Pressure 6-13 6-7 S-II LO× Pump Inlet Conditions 6-14

xi LIST OF ILLUSTRATIONS(CONTINUED)

Figure Page 7-I S-IVB Start Box and Run Requirements - First Burn 7-2 7-2 S-IVB Steady-State Performance - First Burn 7-4 7-3 S-IVB CVSPerformance - Coast Phase 7-6 7-4 S-IVB Ullage Conditions During Repressurization Using 02/H2 Burner 7-9 7-5 S-IVB 02/H2 Burner Thrust and Pressurant Flowrates 7-I0 7-6 S-IVB Start Box and Run Requirements - Second Burn 7-11 7-7 S-IVB Steady-State Performance - Second Burn 7-13 7-8 S-IVB LH2 Ullage Pressure - First Burn and Parking Orbit 7-15 7-9 S-IVB LH2 Ullage Pressure - Second Burn and Translunar Coast 7-16 7-10 S-IVB Fuel Pump Inlet Conditions - First Burn 7-17 7-11 S-IVB Fuel Pump Inlet Conditions - Second Burn 7-18 7-12 S-IVB LOX Tank Ullage Pressure - First Burn and Parking Orbit 7-19 7-13 S-IVB LOXTank Ullage Pressure - Second Burn and Translunar Coast 7-21 7-14 S-IVB LOX Pump Inlet Conditions - First Burn 7-22 7-15 S-IVB LOX Pump Inlet Conditions - Second Burn 7-23 7-16 S-IVB Cold Helium Supply History 7-24 7-17 S-IVB APS Propellants Remaining Versus Range Time, Module No. 1 and Module No. 2 7-26 7-18 S-IVB LOX Dumpand Orbital Safing Sequence 7-28 7-19 S-IVB LOX Dump 7-30 8-I S-IVB Hydraulic System - Second Burn 8-3 8-2 S-IVB Engine Driven Hydraulic PumpSchematic 8-4 9-I Release Rod Force Time History Comparison 9-2 9-2 Longitudinal Load at MaximumBending Moment, CECOand OECO 9-3 MaximumBending Moment Near Max Q 9-4 First Longitudinal Modal Frequencies. During S-IC Powered Flight 9-4

xi i _ LIST OF ILLUSTRATIONS(CONTINUED)

Figure Page 9-5 Longitudinal Acceleration at CMand IU 9-6 9-6 Peak Amplitudes of Vehicle First Longitudinal Mode for AS-504, AS-505, and AS-506 9-7 9-7 Frequency and Amplitude of Longitudinal Oscillations During S-IC Boost 9-8 9-8 Frequency and Amplitude of Longitudinal Oscillations During S-II Stage Boost 9-9 9-9 S-IVB AS-506 and AS-505 17- to 20-Hertz Oscillations Comparison 9-9 9-10 AS-506 S-IVB First Burn MaximumResponse 9-10 9-11 AS-506 and AS-505 First Burn Response 9-10 9-12 Comparison of 45-Hertz Oscillations During AS-505 and AS-506 Second Burn 9-11 9-13 AS-506 Lateral Analysis/Measured Modal Frequency Correlation 9-12 I0-I Trajectory and ST-124M-3 Platform Velocity Comparison (Trajectory Minus Guidance) 10-2 10-2 Trajectory and ST-124M-3 Platform Velocity Comparison Second S-IVB Burn (Trajectory Minus Guidance) 10-3 10-3 AS-506 Characteristic Velocity Error 10-9 I0-4 Attitude CommandsDuring Active Guidance Period I0-II I0-5 Attitude Angles During S-IVB Second Burn 10-12 II-I Pitch Plane Dynamics During S-IC Burn 11-3 11-2 Yaw Plane Dynamics During S-IC Burn 11-4 11-3 Roll Plane Dynamics During S-IC Burn 11-5 11-4 Normal Acceleration During S-IC Burn 11-8 11-5 Pitch and Yaw Plane Wind Velocity and Free-Stream Angles-of-Attack During S-IC Burn 11-9 11-6 Pitch Plane Dynamics During S-II Burn II-I0 11-7 Yaw Plane Dynamics During S-II Burn II-II 11-8 Roll Plane Dynamics During S-II Burn 11-12 11-9 Pitch Plane Dynamics During S-IVB First Burn 11-14

xiii LIST OF ILLUSTRATIONS(CONTINUED)

Figure Page II-I0 Yaw Plane Dynamics During S-IVB First Burn 11-15 II-II Roll Plane Dynamics During S-IVB First Burn 11-16 11-12 Pitch Plane Dynamics During Coast In Parking Orbit II -17 11-13 Pitch Plane Dynamics During S-IVB Second Burn _I-17 11-14 Yaw Plane Dynamics During S-IVB Second Burn 11-18 11-15 Roll Plane Dynamics During S-IVB Second Burn 11-18 11-16 Pitch and Yaw Plane Dynamics Following Translunar Injection 11-19 11-17 Pitch, Yaw and Roll Plane Dynamics During the Maneuver to TD&EAttitude 11-20 11-18 Pitch, Yaw and Roll Plane Dynamics During the Maneuver to Slingshot Attitude 11-21 13-I S-IVB Stage Forward Battery No. 1 Voltage and Current 13-4 13-2 S-IVB Stage Forward Battery No. 2 Voltage and Current 13-4 13-3 S-IVB Stage Aft Battery No. 1 Voltage and Current 13-5 13-4 S-IVB Stage Aft Battery No. 2 Voltage and Current 13-5 13-5 Battery 6DIO Voltage, Current, and Temperature 13-7 13-6 Battery 6D30 Voltage, Current, and Temperature 13-7 13-7 Battery 6D40 Voltage, Current, and Temperature 13-8 16-I S-IC Base Heat Shield Pressure Loading 16-2 16-2 S-II Heat Shield Aft Face Pressure 16-3 16-3 S-II Heat Shield Forward Face Pressure 16-3 16-4 S-II Thrust Cone Pressure 16-4 16-5 S-II Forward Skirt Pressure Loading 16-6 17-I S-IC Base Heat Shield Measurement Locations 17-2 17-2 S-IC Base Heat Shield Total Heating Rate 17-3 17-3 S-IC Base Heat Shield Gas Temperature 17-3 17-4 S-II Heat Shield Aft Face Heat Rate 17-4

xiv LIST OF ILLUSTRATIONS (CONTINUED)

Figure Page 17-5 Heat Shield Aft Radiation Heat Rate 17-5

17-6 S-II Base Gas Temperature 17-6 17-7 Forward Location of Separated Flow 17-7

18-I S-IC Forward Compartment Ambient Temperature 18-3

18-2 S-IC Aft Compartment Temperature 18-4 18-3 Sublimator Performance During Ascent 18-5 18-4 TCS Coolant Control Parameters 18-6

18-5 TCS GN2 Sphere Pressure (D25-601) 18-7

18-6 IU Selected Component Temperatures "18-8

18-7 Inertial Platform GN2 Pressures 18-9 18-8 GBS GN2 Sphere Pressure (DI0-603) 18-10

19-I VHF Telemetry Coverage Summary 19-7 19-2 C-Band Radar Coverage Summary 19-9

19-3 CCS Signal Strength Fluctuations at Hawaii 19-10 19-4 CCS Signal Strength Fluctuations at GDS Wing Station 19-11

19-5 CCS Coverage. Summary 19-13

A-I Scalar Wind Speed at Launch Time of AS-506 A-4 A-2 Wind Direction at Launch Time of AS-506 A-5

A-3 Pitch Wind Speed Component (Wx) at Launch Time of AS-506 A-6

A-4 Yaw Wind Speed Component (Wz) at Launch Time of AS-506 A-7

A-5 Pitch (Sx) and Yaw (S z) Component Wind Shears at Launch Time of AS-506 A-8

A-6 Relative Deviation of Temperature and Density From the PRA-63 Reference Atmosphere, AS-506 A-II

A-7 Relative Deviation of Pressure and Absolute Deviation of the Index of Refraction From the PRA-63 Reference Atmosphere, AS-506 A-12

xv LIST OF TABLES

Table Page

2-I Time Base Summary 2-3

2-2 Significant Event Times Summary 2-4 2-3 Variable Time and Commanded Switch Selector Events 2-10

3-I AS-506 Prelaunch Milestones 3-2

4-I Comparison of Significant Trajectory Events 4-7

4-2 Comparison of Cutoff Events 4-8

4-3 Comparison of Separation Events 4-9

4-4 Stage Impact Location 4-10 4-5 Parking Orbit Insertion Conditions 4-11

4-6 Translunar Injection Conditions 4-16

4-7 Comparison of Slingshot Maneuver Velocity Increment 4-16

4-8 Comparison of Lunar Closest Approach Parameters 4-19 4-9 Heliocentric Orbit Parameters 4-19

5-I S-IC Engine Performance Deviations 5-5

5-2 S-IC Stage Propellant Mass History 5-6 6-I S-II Engine Performance Deviations (ESC +61 Seconds) 6-7

S-II Propellant Mass History 6-10

S-IVB Steady-State Performance - First Burn (STDV +137-Second Time Slice at Standard Altitude Conditions) 7-5

7-2 S-IVB Steady-State Performance - Second Burn (STDV +172-Second Time Slice at Standard Altitude Conditions) 7-14

7-3 S-IVB Stage Propellant Mass History 7-14

7-4 S-IVB APS-Propellant Conditions 7-25 7-5 S-IVB APS Propellant Consumption 7-27

xvi LIST OF TABLES(CONTINUED)

Table Page I0-I Inertial Platform Velocity Comparisons 10-5 I0-2 Guidance Comparisons 10-6 10-3 Guidance Components Differences 10-8 10-4 Start and Stop Times for IGM Guidance Commands I0-I0 10-5 Parking Orbit Insertion Parameters 10-13 10-6 Translunar Injection Parameters 10-13 II-I AS-506 Misalignment and Liftoff Conditions Summary 11-6 11-2 MaximumControl Parameters During S-IC Burn 11-7 11-3 MaximumControl Parameters During S-II Burn 11-13 11-4 MaximumControl Parameters During S-IVB First Burn 11-16 11-5 MaximumControl Parameters During S-IVB Second Burn 11-19 13-I S-IC Stage Battery Power Consumption 13-I 13-2 S-II Stage Battery Power Consumption 13-2 13-3 S-IVB Stage Battery Power Consumption 13-3 13-4 IU Battery Power Consumption 13-6 14-I Commandand Communication System GDSCommands History 14-2 18-I TCS Coolant Flowrates and Pressures 18-6 19-I AS-506 Measurement Summary 19-2 19-2 AS-506 Flight Measurements Waived Prior to Launch 19-3 19-3 AS-506 Measurement Malfunctions 19-4 19-4 AS-506 Launch Vehicle Telemetry Links 19-5 20-I Total Vehicle Mass - S-IC Burn Phase - Kilograms 20-3 20-2 Total Vehicle Mass - S-IC Burn Phase - Pounds Mass 20-4 20-3 Total Vehicle Mass - S-II Burn Phase - Kilograms 20-5 20-4 Total Vehicle Mass - S-II Burn Phase - Pounds Mass 20-6 20-5 Total Vehicle Mass - S-IVB First Burn Phase - Ki I ograms 20-7

xvii LIST OF TABLES(CONTINUED)

Table Page 20-6 Total Vehicle Mass - S-IVB First Burn Phase - Pounds Mass 20-8 20-7 Total Vehicle Mass - S-IVB Second Burn Phase - Kilograms 20-9 20-8 Total Vehicle Mass - S-IVB Second Burn Phase - Pounds Mass 20-10 20-9 Flight Sequence Mass Summary 20-I 1 20-10 Mass Characteristics Comparison 20-13 21-I Mission Objectives Accomplishment Summary 21-I 22-I Summaryof Deviations 22-2 A-I Surface Observations at AS-506 Launch Time A-2 A-2 Solar Radiation at AS-506 Launch Time, Launch Pad 39A A-3 A-3 Systems Used to Measure Upper Air Wind Data for AS-506 A-9 A-4 MaximumWind Speed in High Dynamic Pressure Region for Apollo/Saturn 501 through Apollo/ Saturn 506 Vehicles A-9 A-5 Extreme Wind Shear Values in the High Dynamic Pressure Region for Apollo/Saturn 501 through Apollo/Saturn 506 Vehicles A-IO A-6 Selected Atmospheric Observations for Apollo/ Saturn 501 through Apollo/Saturn 506 Vehicle Launches at Kennedy Space Center, Florida A-13 B-I S-IC Significant Configuration Changes B-2 B-2 S-II Significant Configuration Changes B-2 B-3 S-IVB Significant Configuration Changes B-3 B-4 IU Significant Configuration Changes B-4

xviii ACKNOWLEDGEMENT

This report is published by the Saturn Flight Evaluation Working Group-- composed of representatives of Marshall Space Flight Center, John F. Kennedy Space Center, and MSFC's prime contractors--and in cooperation with the Manned Spacecraft Center. Significant contributions to the evaluation have been made by:

George C. Marshall Space Flight Center Science and Engineering

Central Systems Engineering Aero-Astrodynamics Astrionics Laboratory Computation Laboratory Astronautics Laboratory

Program Management John F. Kennedy Space Center

Manned Spacecraft Center

The Boeing Company

McDonnell Douglas Astronautics Company

International Business Machines Corporation

North American Rockwell/Space Division

North American RockwelJ/Rocketdyne Division

xix ABBREVIATIONS

ACN Ascension DTS Data Transmission System AGC Automatic Gain Control EBW Exploding Bridge Wire ANT Antigua ECO Engine Cutoff AOS Acquisition of Signal ECS Environmental Control System APS Auxiliary Propulsion System EDS Emergency Detection System ARIA Apollo Range Instrument EDT Eastern Daylight Time Aircraft EMR Engine Mixture Ratio ASI Augmented Spark Igniter EPO Earth Parking Orbit AUX Auxiliary ESC Engine Start Command AVP Address Verification Pulse EVA Extra-Vehicular Activity BDA Bermuda FCC Flight Control Computer CCS Commandand Communications System FM/FM Frequency Modulation/ Frequency Modulation CDDT Countdown Demonstration FRT Test Flight Readiness Test GBI Grand Bahama Island. CECO Center Engine Cutoff GBM Grand Bahama CG Center of Gravity GBS ClF Central Information Gas Bearing System Facility GET Ground Elapse Time CM CommandModule GFCV GOXFlow Control Valve CNV Cape Kennedy GDS Goldstone CRO Carnarvon GG Gas Generator CRP Computer Reset Pulse GOX Gaseous Oxygen CSM Commandand Service Module GRR Guidance Reference Release CVS Continuous Vent System GSE Ground Support Equipment CYI Grand Canary Island GSFC Goddard Space Flight Center DDAS Digital Data Acquisition GTK Grand Turk Island System GWM Guam DEE Digital Events Evaluator GYM Guaymas

XX Mixture Ratio HAW Hawaii MR HDA Holddown Arm MSC Manned Spacecraft Center

HFCV Helium Flow Control Valve MSFC Marshall Space Flight Center

HSK Honeysuckle (Canberra) MSFN Manned Space Flight Network Mobile Service Structure I GM Iterative Guidance Mode MSS

IMU Inertial Measurement Unit MTF Mississippi Test Facility IP&C Instrumentation Program M/W Methanol Water and Components NPSP Net Positive Suction Pressure

IU Instrument Unit NPV Non Propulsive Vent KSC Kennedy Space Center NASA National Aeronautics and LCC Launch Control Center Space Administration

LES Launch Escape System OAFPL Overall Fluctuating Pressure Level LET Launch Escape Tower OASPL Overall Sound Pressure Level LH 2 Liquid Hydrogen OAT Overall Test LIEF Launch Information Exchange Facility OCP Orbital Correction Program LM Lunar Module OECO Outboard Engine Cutoff

LOI Lunar Orbit Insertion OIS Operational Intercom System

LOS Loss of Signal OMNI Omni Directional

LOX Liquid Oxygen OT Operational Trajectory LUT Launch Umbilical Tower PAM/ Pulse Amplitude Modulation/ FM/FM Frequency Modulation/ LV Launch Vehicle Frequency Modulation LVDA Launch Vehicle Data PAFB Patrick Air Force Base Adapter PCM Pulse Code Modulation LVDC Launch Vehicle Digital PCM/ Pulse Code Modulation/ Computer FM Frequency Modulation MAD Madrid PDO Principal Detailed Objective MAP Message Acceptance Pulse PMR Programed Mixture Ratio MCC-H Mission Control Center - Houston PRA Patrick Reference Atmosphere

MER Mercury (ship) PSD Power Spectral Density Pad Terminal Connection Room MFCV Modulating Flow Control PTCR Valve PTCS Propellant Tanking Control MILA Merritt Island Launch Area System

MOV Main Oxidizer Valve PU Propellant Utilization

xxi RED Redstone (ship) TMR Triple Modular Redundant RF Radio Frequency TSM Tail Service Mast RMS Root Mean Square TVC Thrust Vector Control RP-I Designation for S-IC Stage USB Unified S-Band Fuel (kerosene) UT Universal Time SA Service Arm VAN Vanguard (ship) SC Spacecraft VHF Very High Frequency SDO Secondary Detailed WHS White Sands Objective SLA Spacecraft LMAdapter SM Service Module SMC Steering Misalignment Correction SPL Sound Pressure Level SPS Service Propulsion System SRSCS Secure Range Safety CommandSystem SS/FM Single Sideband/Frequency Modulation STDV Start Tank Discharge Valve SV Space Vehicle T1 Time Base 1 Tli Time to go in Ist Stage IGM T2i Time to go in 2nd Stage I GM TAN Tananarive TCS Thermal Conditioning Systern TD&E Transposition, Docking and Ejection TEl Transearth Injection TEX Corpus Christi (Texas) TLC Translunar Coast TLI Translunar Injection TM Telemeter, Telemetry

xxi i MISSION PLAN

The AS-506 flight (Apollo II Mission) is the sixth flight of the Apollo/ Saturn V flight test program. The primary objective of the mission is to land astronauts on the lunar surface and return them safely to earth. The crew consists of Neil Armstrong (Mission Commander), Lt. Col. Michael Collins (CommandModule Pilot), and Lt. Col. Edwin Aldrin, Jr. (Lunar Module Pilot).

The AS-506 flight vehicle is composed of the S-IC-6, S-II-6, and S-IVB-6N stages; Instrument Unit (IU)-6; Spacecraft/Lunar Module Adapter (SLA)-I4; and Spacecraft (SC). The SC consists of Commandand Service Module (CSM) -107 and Lunar Module (LM)-5.

Vehicle launch from Complex 39A at Kennedy Space Center (KSC) is along a 90 degree azimuth with a roll to a variable flight azimuth of 72 to 108 degrees measured east of true north. Vehicle mass at S-IC ignition is 2,941,221 kilograms (6,484,282 Ibm). The S-IC stage powered flight is approximately 161 seconds; the S-II stage provides powered flight for approximately 389 seconds. Following S-IVB first burn (approximately 144 seconds duration), the S-IVB/IU/SLA/LM/CSM is inserted into a 183.8 by 186.5 kilometer (99.2 by 100.7 n mi) altitude (referenced to a spherical earth) Earth Parking Orbit (EPO). Vehicle mass at orbit insertion is 135,669 kilograms (229,099 Ibm).

At approximately I0 seconds after EPO insertion, the vehicle is aligned with the local horizontal. Continuous hydrogen venting is initiated shortly after EPO insertion and the Launch Vehicle (LV) and CSMsystems are checked in preparation for the Translunar Injection (TLI) burn. During the second or third revolution in EPO, the S-IVB stage is reignited and burns for approximately 349 seconds. This burn injects the S-IVB/IU/SLA/LM/CSM into a free-return, translunar trajectory.

Approximately 15 minutes after TLI, the vehicle initiates an inertial attitude hold for CSMseparation, docking and LM ejection. Following the attitude freeze, the CSMseparates from the LV and the SLA panels are jettisoned. The CSMthen transposes and docks to the LM. After docking, the CSM/LMis spring ejected from the S-IVB/IU. Following CSM/LMejection, the S-IVB/IU configuration achieves a co-rotational slingshot trajectory by using propulsive venting of hydrogen (LH2), dumping of oxygen (LOX) and by firing the Auxiliary Propulsion ,System (APS) ullage engines. The slingshot trajectory results in a solar orbit for the S-IVB/IU.

xxiii During the 3 day translunar coast, the astronauts perform star-earth landmark sightings, Inertial Measurement Unit (IMU) alignments, general lunar navigation procedures and possibly four midcourse corrections. At approximately 76 hours, a Service Propulsion System (SPS), Lunar Orbit Insertion (LOI) burn of approximately 359 seconds inserts the CSM/LM into a Ill by 315 kilometer (60 by 17O n mi) altitude parking orbit.

After two revolutions in lunar orbit, a 16-second SPS burn circularizes the orbit to Ill kilometers (60 n mi) altitude at 80 hours. The LM is entered by astronauts Armstrong and Aldrin and checkout is accomplished. During the eleventh revolution in orbit at lO0 hours, the LM separates from the CSM and prepares for the lunar descent. The LM descent propulsion system is used to brake the LM into the landing trajectory, approach the landing site and perform the landing at I03 hours.

Following lunar landing, the two astronauts execute a 2.66 hour simultaneous lunar Extra-Vehicular Activity (EVA). After the EVA, the astronauts prepare the ascent propulsion system for lunar ascent. The total lunar stay time for Apollo II is approximately 22 hours.

The CSM performs a plane change approximately 17 hours prior to lunar ascent. At approximately 124.5 hours, the ascent stage inserts the LM into a 16.7 by 83.3 kilometer (9 by 45 n mi) altitude lunar orbit, and rendezvous and docks with the CSM. The astronauts reenter the CSM, jettison the LM and prepare for Transearth Injection (TEl). TEl is accomplished at approximately 135 hours with a 149-second SPS burn. The time and duration of the SPS TEl burn is dependent on an optional astro- naut rest period.

During the 60-hour transearth coast, the astronauts perform navigation procedures, star-earth-moon sightings and possibly three midcourse corrections. The Service Module (SM) separates from the Command Module (CM) 15 minutes prior to reentry. Splashdown occurs in the Pacific Ocean approximately 195 hours after liftoff.

After the recovery operations, a biological quarantine is imposed on the crew and CM. An incubation period of 18 days from splashdown (21 days from lunar ascent) is required for the astronauts. The hardware incubation period is the time required to analyze certain lunar samples.

xxiv FLIGHT TEST SUMMARY

The fourth manned Saturn V Apollo space vehicle, AS-506 (Apollo II (Mission) was launched at 09:32:00 Eastern Daylight Time (EDT) on July 16, 1969 from Kennedy Space Center (KSC), Complex 39, Pad A. This sixth launch of the Saturn V/Apollo successfully performed the three principal detailed objectives mandatory for successful accomplishment of the primary mission objective which was to perform a lunar landing and return. The secondary detailed objective was also successfully accomplished.

The launch countdown was completed without any unscheduled countdown holds. Ground system performance was satisfactory. Damageto the pad, Launch Umbilical Tower (LUT) and support equipment was minor. The trajectory parameters of AS-506 from launch to Translunar Injection (TLI) were close to nominal. The vehicle was launched on an azimuth 90 degrees east of north. A roll maneuver was initiated at 13.2 seconds that placed the vehicle on a flight azimuth of 72.058 degrees east of north. The space-fixed velocity at S-IC Outboard Engine Cutoff (OECO) was 8.5 m/s (27.9 ft/s) greater than nominal. The space-fixed velocity at S-II OECOwas 22.8 m/s (74.8 ft/s) less than nominal. The space-fixed velocity at S-IVB first guidance cutoff was 0.2 m/s (0.6 ft/s) less than nominal. The altitude at S-IVB first guidance cutoff was 0.2 kilometer (0.I n mi) lower than nominal and the surface range was 1.7 kilometer (I.0 n mi) less than nominal. The space-fixed velocity at parking orbit insertion was equal to nominal. The apogee and perigee were 0.5 kilometer (0.3 n mi) and 0.6 kilometer (0.3 n mi) less than nominal, respectively. The parameters at TLI were also close to nominal. The space- fixed velocity was 3.2 m/s (10.5 ft/s) greater than nominal, the altitude was 3.1 kilometers (1.6 n mi) less than nominal and C3 was 16,877 m2/s2 (181,663 ft2/s 2) greater than nominal. Following Lunar Module (LM) extraction, the vehicle maneuvered to a slingshot attitude frozen relative to local horizontal. The retrograde velocity change necessary to achieve S-IVB/IU lunar slingshot maneuver was accomplished by a LOX dump, Auxiliary Propulsion System (APS) burn, and LH2 vent. The S-IVB/IU closest approach of 3379 kilometers (1825 n mi) above the lunar surface occurred at 78.7 hours into the mission.

All S-IC propulsion systems performed satisfactorily and the propulsion performance level was very close to nominal. Stage site thrust (averaged from liftoff to OECO)was 0.62 percent lower than predicted. Total propellant consumption rate was 0.40 percent lower than predicted with the total consumed Mixture Ratio (MR) 0.I0 percent lower than predicted.

xxv Specific impulse was 0.16 percent lower than predicted. Total propellant consumption from Holddown Arm (HDA) release to OECOwas low by 1.12 percent. Center Engine Cutoff (CECO)was commandedby the IU as planned. OECO,initiated by the LOX low level sensors, occurred 0.55 second later than predicted. The S-II propulsion system performed satisfactorily throughout the flight. The S-II stage operation time of 385.18 seconds was 4.0 seconds shorter than predicted. Early CECOsuccessfully avoided high amplitude low frequency oscillations experienced on the AS-503 and AS-504 flights. Total stage thrust at 61 seconds after S-II Engine Start Command(ESC) was 0.20 percent below predicted. Total propellant flowrate (including pressurization flow) was 0.13 percent below predicted and vehicle specific impulse was 0.07 percent below predicted at this time slice. Stage propellant MR was 0.36 percent above predicted. The engine servicing system Ground Support Equipment (GSE) performed satisfactorily except that the engine No. 1 start tank pressure was 2.8 N/cm2 (4 psi) below redline at prelaunch commit (-33 seconds). All start tank pressures and temperatures were well within requirements at S-II ESC,

The J-2 engine operated satisfactorily throughout the operational phase of S-IVB first and second burn. Shutdowns for both burns were normal. S-IVB first burn duration was 147.1 seconds which was 3.4 seconds more than predicted. The engine performance during first burn, as determined from standard altitude reconstruction analysis, deviated from the predicted by +0.20 percent for thrust and +0.05 percent for specific impulse. The S-IVB stage first burn Engine Cutoff (ECO) was initiated by the Launch Vehicle Digital Computer (LVDC) at 699.34 seconds. The Continuous Vent System (CVS) adequately regulated LH2 tank ullage pressure during orbit, and the Oxygen/Hydrogen (02/H 2) burner satisfactorily achieved LH2 and LOX tank repressurization for restart. Engine restart conditions were within specified limits. The restart at full open Propellant Utilization (PU) valve position was successful. S-IVB second burn duration was 346.9 seconds which was 1.7 seconds less than predicted. The engine performance during second burn, as determined from the standard altitude reconstruction analysis, deviated from the predicted by -0.56 percent for thrust and +0.05 percent for specific impulse. Subsequent to second burn, the stage propellant tanks were safed satisfactorily.

The stage hydraulic systems performed satisfactorily on the S-IC, S-II, and first burn and coast phase of the S-IVB stage. During this period all parameters were within specification limits. Just after stage reignition the S-IVB hydraulic system pressure exceeded the upper limit by 0.6 percent. At 202 seconds into the burn, a step decrease in system pressure to a nominal operating level occurred and the pressure remained at this level for the remainder of the burn. The pumpmanufacturer does not consider this condition to indicate impending malfunction of the engine driven pump.

xxvi The structural loads and dynamic environments experienced by the AS-506 launch vehicle were well within the vehicle structural capability. The longitudinal loads experienced during flight were nominal. The maximum bending moment condition, 3.75 x 106 N-m (33.2 x 106 Ibf-in.), was experienced at 91.5 seconds and was lower than that experienced on any previous flight. Low level first mode longitudinal oscillations similar to those of previous flights were evident during each stage burn but caused no problems.

The navigation and guidance system performed satisfactorily. The parking orbit and TLI parameters were well within tolerance. The S-IVB LOX dump, LH2 vent and APS ullage burn resulted in a heliocentric orbit of the S-IVB/IU as planned. The actual S-II Engine Mixture Ratio (EMR) shift occurred approximately 9.5 seconds later than indicated by the final stage propulsion prediction. About 4 seconds of this deviation was attributed to the change in LVDCnominal characteristic velocity presetting predictions and variation in actual from predicted flight performance. Approximately 5.5 seconds of the deviation is attributed to improper scaling in the flight program calculation of characteristic velocity. The LVDC, the Launch Vehicle Data Adapter (LVDA), and the ST-124M-3 inertial platform functioned satisfactorily. The platform- measured crossrange velocity (_) exhibited a negative shift of approximately 1.8 m/s (5.9 ft/s) at 3.3 seconds after liftoff. The probable cause was the Y accelerometer head momentarily contacting an internal mechanical stop. This had negligible effect on launch vehicle performance.

The AS-506 Flight Control Computer (FCC), Thrust Vector Control (TVC) and APS satisfied all requirements for vehicle attitude control during the flight. All maneuvers were properly accomplished. All separations occurred as expected without producing significant attitude deviations.

The AS-506 launch vehicle electrical systems performed satisfactorily throughout all phases of flight. Performance of the Secure Range Safety CommandSystems (SRSCS)was nominal on all powered stages. The SRSCS was properly safed by ground commandfrom Bermuda (BDA). Performance of the Commandand Communications System (CCS) was satisfactory except for the Radio Frequency (RF) problem noted. The Emergency Detection System (EDS) performance was nominal with no abort limits exceeded.

Vehicle base pressure environments were generally in good agreement with postflight predictions and compared well with previous flight data. There was no instrumentation provided on the AS-506 vehicle which would permit a direct evaluation of the surface and compartment pressure environments. The one ambient pressure measurement located in the S-II forward skirt was used to calculate the pressure loading acting on that area, and indicated good agreement with postflight predictions and previous flight data.

xxvii Base thermal environments were similar to those experienced on earlier flights with the exception that S-II heat shield aft radiation heating rates were approximately 20 percent higher than the maximumvalues measured during previous flights. Aerodynamic heating environments were not measured on AS-506,

The Environmental Control System (ECS) performed satisfactorily. The IU ECS coolant temperatures, pressure, and flowrates were continuously maintained within required ranges and design limits. One deviation from specification was observed. The inertial platform gas bearing differential2Pressure drifted above the 10.7 N/cm2 (15.5 psid) maximum to 11.2 N/cm (16.3 psid). This condition has occurred on previous flights and caused no detrimental effect on the missions.

All elements of the data system performed satisfactorily except for a problem with the CCSdownlink during translunar coast. Measurement performance was excellent as evidenced by 99.9 percent reliability. This is the highest reliability attained on any Saturn V flight. Tele- metry performance was nominal, with the exception of a minor calibration deviation. Very High Frequency (VHF) telemetry Radio Frequency (RF) propagation was generally good, though the usual problems due to flame effects and staging were experienced. VHF data were received to 17,800 seconds (04:56:40). Commandsystems RF performance for both the SRSCSand CCSwas nominal except for the CCSdownlink problem noted. Goldstone (GDS) reported receiving CCSsignals to 35,779 seconds (9:56:19). Good tracking data were received from the C-Band radar, with Patrick Air Force Base (PAFB) indicating final LOS at 42,912 seconds (11:55:12). The 75 ground engineering cameras provided good data during the launch.

xxviii SECTION1

INTRODUCTION

I.I PURPOSE

This report provides the National Aeronautics and Space Administration (NASA) Headquarters, and other interested agencies, with the launch vehicle evaluation results of the AS-506 flight test. The basic objective of flight evaluation is to acquire, reduce, analyze, evaluate and report" on flight test data to the extent required to assure future mission success and vehicle reliability. To accomplish this objective, actual flight failures, anomalies and deviations must be identified, their causes accurately determined, and complete information made available so that corrective action can be accomplished within the established flight schedule.

1.2 SCOPE

This report presents the results of the early engineering flight evaluation of the AS-506 launch vehicle. The contents are centered on the performance evaluation of the major launch vehicle systems, with special emphasis on the deviations. Summaries of launch operations and spacecraft performance are included for completeness. The official George C. Marshall Space Flight Center (MSFC) position at this time is represented by this report. It wili not be followed by a similar report unless continued analysis or new information should prove the conclusions presented herein to be significantly incorrect. Final stage evaluation reports will, however, be published by the stage contractors. Reports covering major subjects and special subjects will be published as required.

1-I/I-2

SECTION2

EVENTTIMES

2.1 SUMMARYOF EVENTS

Range zero time, the basic time reference for this report, is 9:32:00 Eastern Daylight Time (EDT) (13:32:00 Universal Time [UT]). This time is based on the nearest second prior to S-IC tail plug disconnect which occurred at 9:32:00.6 EDT. Range time is calculated as the elasped time from range zero time and, unless otherwise noted, is the time used throughout this report. The actual and predicted range times are adjusted to ground telemetry received times. The Time-From-Base times are presented as vehicle times. Figure 2-I shows the time delay of ground telemetry received time versus Launch Vehicle Digital Computer (LVDC) time and indicates the magnitude and sign of corrections applied to correlate range time and vehicle time in Tables 2-I, 2-2 and 2-3.

Guidance Reference Release (GRR) occurred at -16.97 seconds and start of Time Base 1 (T I) occurred at 0.63 seconds. GRRwas established by the Digital Events Evaluator (DEE-6) and T1 was initiated at detection of liftoff signal provided by de-energizing the liftoff relay in the Instru- ment Unit (IU) at IU umbilical disconnect.

Range time for each time base used in the flight sequence program and the signal for initiating each time base are presented in Table 2-I.

Start of T2 was within nominal expectations for this event. Start of T3, T4 and T5 were initiated approximately 0.6 second late and 3.5 and 0.I seconds early, respectively, due to Variations in the stage burn times. These variations are discussed in Sections 5, 6 and 7 of this document. Start of T6, which was initiated by the LVDCupon solving the restart equation, was 0.9 second later than predicted. Start of T7 was 1.0 second earlier than predicted. T , which was initiated by the receipt of a ground command, was starte_ 63.2 seconds later than the predicted time.

A summary of significant events for AS-506 is given in Table 2-2. Since not all events listed in Table 2-2 are IU commandedswitch selector functions, deviations are not to be construed as failures to meet specified switch selector tolerances. The events in Table 2-2 associated with guidance, navigation, and control have been identified as being accurate to within a major computation cycle.

2-I 320

280

240 4 F

/ ,/ . 200 /

_ 160

Z 120 / i-- / 80 / J 40 I ,...... -- I

0 S_

0 2 4 6 8 I0 12 14 16 18 20 22 RANGE TIME, I000 SECONDS I I I I I i I 0:00:00 I:00:00 2:00:00 3:00:00 4:00:00 5:00:00 6:00:00

RANGE TIME, HOURS:MINUTES:SECONDS

Figure 2-I. Telemetry Time Delay

The predicted times for establishing actual minus predicted times in Table 2-2 have been taken from 40M33626B, "Interface Control Document Definition of Saturn SA-506 Flight Sequence Program", and from the "AS-506 G Mission Launch Vehicle Operational Trajectory", dated July 14, 1969.

2.2 VARIABLE TIME AND COMMANDED SWITCH SELECTOR EVENTS

Table 2-3 lists the switch selector events which were issued during the flight but were not programed for specific times. The range times are adjusted to ground telemetry received times. The water coolant valve open and close switch selector commands were issued based on the condition of two thermal switches in the Environmental Control System (ECS). The outputs of these switches were sampled once every 300 seconds, beginning at 180 seconds, and a switch selector command was issued to open or close the water valve. The valve was opened if the sensed temperature was too high and the valve was closed if the temperature was too low.

2-2 Table 2-I. Time Base Summary

RANGE TIME TIME BASE SEC SIGNAL START (HR:MIN:SEC)

T o -16.97 Guidance Reference Release

T1 0.63 IU Umbilical Disconnect Sensed by LVDC

T 2 135.27 S-IC CECO Sensed by LVDC

T 3 161.66 S-IC OECO Sensed by LVDC

T4 548.24 S-II OECO Sensed by LVDC

T 5 699.57 S-IVB ECO (Velocity) Sensed by LVDC

T6 9278.24 Restart Equation Solution (2:34:38.24)

T 7 10,203.33 S-IVB ECO (Velocity) Sensed (2:50:03.33) by LVDC

T8 17,467.64 Enabled by Ground Command (4:51:07.64)

Table 2-3 also contains the special sequence of switch selector events which were programed to be initiated by telemetry station acquisition and included the following calibration sequence:

Function Stage Time (Sec)

Telemetry Calibrator IU Acquisition +60.0 In-Flight Calibrate ON

TM Calibrate ON S-IVB Acquisition +60.4

TM Calibrate OFF S-IVB Acquisition +61.4

Telemetry Calibrator IU Acquisition +65.0 In-Flight Calibrator OFF

2-3 Table 2-2. Significant Event Times Summary

RANGE TIME TIME FROM BASE EVENI DESCRIPTION ACTUAL ACT-PRED ACTUAL ACT-PRED SEC SEC SEC SEC

I GUIDANCE REFERENCE RELEASE -17.0 0.0 -17.6 0.0 (GRR)

2 S-IC ENGINE START SEQUENCE -8.9 0.0 -9.5 0.0 COMMAND (GROUND)

3 S-IC ENGINE NO.I START -6. I 0.0 -6.8 0.0

4 S-IC ENGINE NO.2 START -5.9 0-0 -6.5 0.0

5 S-IC ENGINE NO.3 START -6.1 0.0 -6.7 0.0

6 S-IC ENGINE NO.4 START -6.0 0.0 -6.6 0.0

7 S-IC ENGINE NO.5 START -6.4 0.0 -7. I 0.0

B ALL S-IC ENGINES THRUST OK -1.6 -0.1 -2.2 -0.I

9 RANGE ZERO 0.0 -0.6

I0 ALL HOLDDOWN ARMS RELEASED 0.3 0.0 -0.3 0.1 (FIRST MOTICN)

11 IU UMBILICAL DISCONNECT, START 0.6 -0.I 0.0 0.0 OF TIME BASE I (T[)

12 BEGIN TOWER CLEARANCE YAW 1.7 0.0 1.0 0.0 MANEUVER*

13 END YAW MANEUVER* 9.? -1.O 9.0 -1.O

14 BEGIN PITCH AND ROLL MANEUVER* 13.2 -0.6 12.6 -0.5

15 S-IC OUTBOARD ENGINE CANT 20.6 -0.I 20.0 0.0

16 END ROLL MANEUVER * 31.1 -0.7 30.5 -0.6

17 MACH 1 66.3 0.7 65.7 0.7

18 MAXIMUM DYNAMIC PRESSURE 83.0 1.7 82.4 1.8 (MAX Q)

19 S-IC CENTER ENGINE CUTOFF 135.20 -0.08 134.56 -0.06 (CECO)

20 START OF TIME BASE 2 (T2) 135.3 0.0 0.0 0.0

21 END PITCH MANEUVER (TILT 160.0 -0.8 24.7 -0.8 ARREST)*

22 S-IC OUTBOARD ENGINE CUTOFF 161.63 0.55 26.36 0.59 (OECO)

23 START GF TIME BASE 3 (13) 161.? 0.6 0.0 0.0

24 START S-If LH2 TANK HIGH 161.7 0.5 0.I 0.0 PRESSURE VENT MODE

*Time is accurate to major computation cycle dependent upon length of computation cycles.

2-4 Table 2-2. Significant Event Times Summary (Continued)

RANGE TIME TIME FROM BASE EVENT DESCRIPTION ACTUAL ACT-PRED ACTUAL ACT-PRED SEC SEC SEC SEC

25 S-II LH2 RECIRCULAIION PUMPS 161.8 0.5 0.2 0.0 OFF

26 S-If ULLAGE MOTOR IGNITION 0.5 0.5 0.0

27 S-ICIS-II SEPARATION COMMAND 0.5 0.7 0.0 TO FIRE SEPARATION DEVICES AND RETRO MOTORS

28 S-IT ENGINE START COMMAND 163.0 0.5 1.4 0.0 (ESC)

29 S-IT ENGINE SOLENOID ACTIVAT- 164.0 0.5 2.4 0.0 ION (AVERAGE OF FIVE)

30 S-II ULLAGE MOTOR BURN TIME 166.1 0.4 4.4 -0.2 TERMINATION (THRUST REACHES 75%)

31 S-II MAINSTAGE 166.2 0.7 4.6 0.2

32 S-IT CHILLDOWN VALVES CLOSE 168.0 0.5 6.4 0.0

33 ACTIVATE S-IT PU SYSTEM 168.5 0.5 6.9 0.0

34 S-II SECOND PLANE SEPARATION 192.3 0.5 30.7 0.0 COMMAND {JETTISON S-IT AFT INTERSTAGE)

35 LAUNCH ESCAPE TOWER (LET) 197.9 0.4 36.2 -0°2 JETTISON

36 ITERATIVE GUIDANCE MODE (IGM) 204.1 1.5 42.4 0.9 PHASE I INITIATED*

3T S-IT LOX STEP PRESSURIZATION 261.6 IO0.O 0.0

38 S-II CENTER ENGINE CUTOFF 460.6 299.0 0°0 (CECO)

39 S-II LH2 STEP PRESSURIZATION 461.6 0.5 300°0 O.O

40 GUIDANCE SENSED TIME TO BEGIN 494.8 6.0 333.2 5.5 EMR SHIFT (IGM PHASE 2 INI- TIATED & START OF ARTIFI- CIAL TAU MODE)*

41 S-IT LOW ENGINE MIXTURE RATIO 498.0 336.3 (EMR) SHIFT (ACTUAL)

42 END OF ARTIFICIAL TAU MODE * 504.2 342.5

43 S-IT OUTBOARD ENGINE CUTOFF 548.22 386.56 {OECO)

44 S-I[ ENGINE CUTOFF INTERRUPTt 548.2 -3.5 0.0 0°0 START OF TIME BASE 4 (I4) {START OF IGM PHASE 3)

*Time is accurate to major computation cycle dependent upon length of computation cycles.

2-5 Table 2-2 Significant Event Times Summary (Continued)

RANGE TIME TIME FROM BASE EVENT DESCRIPTICN ACTUAL ACT-PRED ACTUAL ACT-PRED SEC SEC SEC SEC

45 S-IVB ULLAGE MOTOR IGNITION 0.7 0.0

46 S-II/S-IVB SEPARATION COMMAND 0.8 0.0 TO FIRE SEPARATION DEVICES AND RETRO MOTORS

47 S-IVB ENGINE START COMMAND 549.2 -3.4 1.0 0.0 (FIRST ESC)

48 FUEL CHILLDDWN PUMP CFF 550.4 -3.5 2.2 0.0

49 S-IVB IGNITION (STDV OPEN) 552.2 -3.5 4.0 0.0

50 S-IVB PAINSTAGE 554.7 -3.5 6.5 0.0

51 START UF ARTIFICIAL TAU MODE* 555.6 -5.7 7.3 -2.3

52 S-IVB ULLAGE CASE JETTISON 561.0 -3.4 12.8 O.O

53 END OF ARTIFIEIAL TAU MODE * 562.4 -8.9 14.2 -5.4

54 BEGIN TERMINAL GUIDANCE* 665.2 -0.3 I16.9 3.0

55 END IGM PHASE 3 * 691.6 -0.2 143.4 3.3

56 _EGIN CHI FREEZE * 691.6 -0.2 143.4 3.3

57 S-IVB VELOCITY CUTOFF COMMAND 699.34 -0.15 -0.23 -0;03 (FIRST GUIDANCE CUTOFF) (FIRST ECO)

58 S-IVB ENGINE CUTOFF INTERRUPT, 699.6 -0.1 0.0 0°0 START OF TIME BASE 5 (I5)

59 S-IVB APS ULLAGE ENGINE NO. L 699.8 -0.2 0.3 0.0 IGNITION CCMMAND

60 S-IVB APS ULLAGE ENGINE NO. 2 699.9 -0.2 0.4 O.O IGNITION COMMAND

61 LOX TANK PRESSURIZATICN OFF 700.7 -0.2 1.2 0.0

62 PARKING ORBIT INSERTION 709.34 -0.15 9.77 -0.04

63 BEGIN MANEUVER TO LOCAL 719.3 -0.5 19.8 -0.3 HORIZONTAL ATTITUDE *

64 S-IVB LH2 CONTINUOUS VENT 758.5 -0.2 59.0 0.0 SYSTEM {CVS) ON

65 S-IVB APS ULLAGE ENGINE NO. I 786.5 -0.2 87.0 0.0 CUTOFF COMMAND

66 S-IVB APS ULLAGE ENGINE NO. 2 786.6 -0.2 87.1 0.0 CUTOFF COMMAND

67 FIRST ORBITAL NAVIGATION 801.1 0.8 101.5 0.9 CALCULATIONS*

*Time is accurate to major computation cycle dependent upon length of computation cycles

2-6 Table 2-2. Significant Event Times Summary (Continued)

RANGE TIME TIME FROM BA_E EVENT DESCRIPTION ACTUAL ACT-PRED ACTUAL AC T-PRED SEC SEC SEC SEC

68 BEGIN S-IVB RESTART PREPARA- 9278.2 0.9 O.O 0.0 TIONS, START OF TIME BASE 6 (T6)

69 S-IVB 02/H2 BURNER LH2 ON 9319.5 0.9 41.3 0.0

70 S-IVB 02/H2 BURNER EXCITERS ON 9319.8 0.9 41.6 O.O

71 S-IVB 02/H2 BURNER LOX ON 9320.2 0.9 42.0 0.0 (HELIUM HEATER ON)

72 S-IVB LH2 VENT OFF (CVS OFF) 9320.4 0.9 62.2 0.0

73 S-IVB LH2 REPRESSURIZATION 9326.3 0.9 48.1 0.0 CONTROL VALVE ON

74 S-IVB LOX REPRESSURIZATION 9326.5 0.9 48.3 0.0 CONTROL VALVE ON

75 S-IVB AUX HYDRAULIC PUMP 9497.2 0.9 21q.0 0.0 FLIGHT MODE ON

76 S-IVB LOX CHILLDOWN PUMP ON 9527.2 0.9 249.0 0.0

77 S-IVB LH2 CHILLDOWN PUMP ON 9532.2 0.9 254.0 0.0

78 S-IVB PREVALVES CLOSED 9537.2 0.9 259.0 0.0

79 S-IVB PU MIXTURE RATIO 4.5 ON 9728.3 0.9 450.I 0.0

80 S-lVB APS ULLAGE ENGINE NO. I 9774.5 0.9 496.3 O.O IGNITION COMMAND

81 S-IVB APS ULLAGE ENGINE NO. 2 9774.6 0.9 496.4 0.O IGNITION COMMAND

82 S-IVB 02/H2 BURNER LH2 OFF 9775.0 0.9 496.8 0.0 (HELIUM HEATER OFF)

83 S-IVB 021H2 BURNER L0X OFF 9779.5 0.9 501.3 0.0

84 S-IVB LH2 CHILLDOWN PUMP OFF 9847.6 0.9 569.4 0.0

85 S-IVB LOX CHILLDOWN PUMP OFF 9847.8 0.9 569.6 O.O

86 S-IVB ENGINE RESTART COMMAND 9848.2 0.9 570.0 o.o (FUEL LEAD INITIATION) (SECCND ESC)

87 S-IVB APS ULLAGE ENGINE NO. I 9851.2 0.9 573.0 O.O CUTOFF COMMAND

88 S-IVB APS ULLAGE ENGINE NO. 2 9851.3 0.9 573.1 0.0 CUTOFF COMMAND

89 S-IVB SECOND IGNITION (STDV R856.2 0.7 578.0 -0.2 OPEN)

2-7 Table 2-2. Significant Event Times Summary (Continued)

RANGE TIME TIME FROM BASE EVENT DESCRIPTION ACTUAL ACT-PRED ACTUAL ACT-PRED SEC SEC SEC SEC

90 S-IVB MAINSTAGE 9858.7 580.5 -0.2

91 ENGINE MIXTURE RATIO (EMR) 9974.4 696.2 -1.5 SHIFT

92 S-IV8 LH2 STEP PRESSURIZATION 10128.2 0.9 850.0 0.0 (SECOND BURN RELAY OFF)

93 BEGIN TERMINAL GUIDANCE* 10174.5 -0.5 896.3 -I .4

94 BEGIN CHI FREEZE * 10201.9 0.6 923.7 -0.3

95 S-IVB SECOND GUIDANCE CUTOFF 10203.07 -I.0 -0.26 -0.06 COMMAND (SECOND ECO)

96 S-IVB ENGINE CUTOFF INTERRUPTv 10203.3 -I.0 0.0 0.0 START OF TIME BASE 7

97 LH2 VENT ON COMMAND 10203.8 -I.0 0.0

98 IRANSLUNAR INJECTION 10213.07 -l.O -0.05

99 BEGIN MANEUVER TO LOCAL 10223.0 -2.8 19.7 -I.8 HORIZONTAL ATTITUDE *

I00 FIRST ORBITAL NAVIGATION 10223.9 -1.9 20.6 -0.9 CALCULATIONS*

101 LH2 VENT OFF COMMAND 11103.1 0.0

102 BEGIN MANEUVER TO TRANSPOSI- 11103.9 0.6 TION AND DOCKING ATTITUDE {TD&E)*

103 CSM SEPARATICN 11723.0 18.7 1519.7 19.7

104 CSM DOCK 12243.7 109.3 2040.4 llO.k

105 SC/LV FINAL SEPARATION 15423.0 418.6 5219.7 419.7

106 START OF TIME BASE 8 (18) 17467.6 63.2 O.O 0.0

107 INITIATE MANEUVER TO SLINGSHOT 17467.6 63.3 0.0 0.0 ATTITUDE *

I08 S-IV8 LH2 VENT ON ICVS ON) 17468.0 63.3 0.4 0.0

109 BEGIN LOX DUMP 18187.6 63.3 720.0 0.0

110 END LOX DUMP 18295.8 63.3 828.2 0.0

111 H2 NCNPROPULSIVE VENT {NPV) ON 19500.6 63.1 2032.9 -0.1

112 S-IVB APS ULLAGE ENGINE NO. I 20267.6 63.3 2800.0 0.0 IGNITION COMMAND

113 S-IVB APS ULLAGE ENGINE NO. 2 20267.8 63.5 2800.2 0.0 IGNITION COMMAND

*Time is accurate to major computation cycle dependent upon length of computation cycles.

2-8 Table 2-2. Significant Event Times Summary (Continued)

RANG E TIME TIME FRCM BASE EVENT DESCRIPTICN ACTUAL ACT-PRED ACTUAL ACT-PRED SEC SEC SEC SEC

Z14 S-IVB APS ULLAGE ENGINE NC. I 20547.6 63.3 3080.0 0.0 CUTOFF COMMAND

115 S-IVB APS ULLAGE ENGINE NO. 2 20547.8 63.5 3080.2 0.O CUTOFF COMMAND

II_ INITIATE MANEUVER TC COMMUNI- 20568.8 64.5 3101.1 I.I CATICNS ATTITUDE

2-9 Table 2-3. Variable Time and CommandedSwitch Selector Events

RANGE TIME TIME FROM BASE FUNCTION STAGE REMARKS (SEC) (SEC)

Water Coolant Valve Open IU 181.0 T3 +19.4 LVDC Function

High (5.5) Engine Mixture S-II 495.8 T3 +334.1 LVDC Function Ratio Off

Low (4.5) Engine Mixture S-II 496.0 T3 +334.3 LVDC Function Ratio On

Water Coolant Valve Closed IU 783.2 T5 +83.7 LVDC Function

TM Calibrate On S-IVB 1057.7 T5 +358.1 CYI Rev 1

TM Calibrate Off S-IVB 1058.7 T5 +359.1 CYI Rev 1

Water Coolant Valve Close IU 3186.9 T5 +2487.4 LVDC Function

Telemetry Calibrator IU 3201.3 T5 +2501.8 CRO Rev 1 Inflight Calibrate On

TM Calibrate On S-lVB 3201.7 T5 +2502.2 CRO Rev 1

TM Calibrate Off S-IVB 3202.7 T5 +2503.2 CRO Rev 1

Telemetry Calibrator IU 3206.3 T5 +2506.8 CRO Rev 1 Inflight Calibrate Off

TM Calibrate Off S-IVB 3642.6 T5 +2943.1 HSK Rev 1

Telemetry Calibrator IU 3646.2 T5 +2946,7 HSK Rev 1 Inflight Calibrate Off

Telemetry Calibrator IU 5369.2 T5 +4669.7 GYM Rev 1 Inflight Calibrate On

TM Calibrate On S-IVB 5369.6 T5 +4670.1 GYM Rev 1

TM Calibrate Off S-IVB 5370.6 T5 +4671.1 GYM Rev 1

Telemetry Calibrator IU 5374.2 T5 +4674.7 GYM Rev 1 Inflight Calibrate Off

Telemetry Calibrator IU 7825.2 T5 +7125.7 TAN Rev 1 Inflight Calibrate On

TM Calibrate On S-IVB 7825.6 T5 +7126.1 TAN Rev 1

TM Calibrate Off S-IVB 7826.6 T5 +7127.1 TAN Rev 1

Telemetry Calibrator IU 7830.2 T5 +7130.7 TAN Rev 1 Inflight Calibrate Off

2-10 Table 2-3. Variable Time and CommandedSwitch Selector Events (Continued)

RANGE TIME TIME FROM BASE FUNCTION STAGE (SEC) (SEC) REMARKS

Telemetry Calibrator IU 8793.3 T5 +8093.8 CRO Rev 2 Inflight Calibrate On

TM Calibrate On S-IVB 8793.7 T5 +8094.2 CRO Rev 2

TM Calibrate Off S-IVB 8794.7 T5 +8095.2 CRO Rev 2

Telemetry Calibrator IU 8798.3 T5 +8098.8 CRO Rev 2 Inflight Calibrate Off

Telemetry Calibrator IU 9678.4 T6 +400.2 ARIA No. 3 Rev 2 Inflight Calibrate On

TM Calibrate On S-IVB 9678.6 T6 +400.4 ARIA No. 3 Rev 2

TM Calibrate Off S-IVB 9679.6 T6 +401.4 ARIA No. 3 Rev 2

Telemetry Calibrator IU 9683.4 T6 +405.2 ARIA No. 3 Rev 2 Inflight Calibrate Off

Water Coolant Valve Open IU 13,409.5 T7 +3206.1 LVDC Function

Water Coolant Valve Close IU 13,803.7 T7 +3507.1 LVDC Function

Water Coolant Valve Open IU 17,319.4 T7 +7116.0 LVDC Function

Passivation Enable S-IVB 18,503.5 T8 +1035.8 OCS Command

Engine He Control Valve S-IVB 18,505.0 T8 +1037.3 CCS Command Open On

TM Calibrate On IU 27,371.9 T8 +9904.0 AcquisitionGYM during TLCby TM Calibrate Off IU 27,372.0 T8 +9904.1

Antenna switching times are not available due to noisy telemetry.

2-11/2-12

SECTION3

LAUNCHOPERATIONS

3.1 SUMMARY

The ground systems supporting the AS-506/ApolIo II countdown and launch performed exceptionally well. Several systems experienced component failures and malfunctions which required corrective actions, but all repairs were accomplished in parallel with the scheduled countdown operations. No unscheduled holds were incurred. Propellant tanking was accomplished satisfactorily. The start of S-II LH2 loading was delayed 25 minutes due to a communications problem in the Pad Terminal Connection Room(PTCR). However, this delay time was recovered during the scheduled hold at -3 hours 30 minutes. Launch occurred at 09:32:00 Eastern Daylight Time (EDT), July 16, 1969, from Pad 39A of the Saturn Complex. Damage to the pad, Launch Umbilical Tower (LUT) and support equipment was minor. 3.2 PRELAUNCHMILESTONES

A chronological summary of events and preparations leading to the launch of AS-506 is contained in Table 3-I.

3.3 COUNTDOWNEVENTS

The AS-506/Apollo II terminal countdown was picked up at -28 hours on July 14, 1969 at 17:00:00 EDT. Scheduled holds of II hours duration at -9 hours in the count, and 1 hour 32 minutes duration at -3 hours 30 minutes, were the only holds initiated. The start of S-II LH2 loading was delayed 25 minutes due to a communications problem in the PTCR. How- ever, this time was recovered during the hold at -3 hours 30 minutes and Space Vehicle (SV) activities were on schedule when the countdown resumed. Launch occurred at 09:32:00 EDT, July 16, 1969, from Pad 39A of the Saturn Complex. 3.4 PROPELLANTLOADING

3.4.1 RP-I Loading

The RP-I system supported the launch countdown satisfactorily. At approximately -21 hours the Propellant Tanking Control System (PTCS) RP-I level indication from the propellant monitor program display became erratic. The problem was traced to a noisy RP-I loading electronics unit. Since

3-I Table 3-I. AS-506 Prelaunch Milestones

DATE ACTIVITY OREVENT

January 8, 1969 LM-5 Ascent Stage Arrival January I0, 1969 SLA-14 Arrival January 12, 1969 LM-5 Descent Stage Arrival January 15, 1969 CSMQuads Arrival January 19, 1969 S-IVB-6N Stage Arrival January 22, 1969 CSM107 Arrival February 6, 1969 S-II-6 Stage Arrival February 20, 1969 S-IC-6 Stage Arrival February 21, 1969 S-IC Erection February 27, 1969 IU-6 Arrival March 4, 1969 S-II Erection March 5, 1969 S-IVB and IU Erections March 18, 1969 CSMAltitude Test with Prime Crew March 21, 1969 LM Altitude Test with Prime Crew March 27, 1969 Launch Vehicle (LV) Propellant Dispersion/ Malfunction Overall Test (OAT) April 14, 1969 Spacecraft (SC) Erection May 5, 1969 Space Vehicle (SV) Electrical Mate May 14, 1969 SV OATNo. 1 (Plugs In) May 20, 1969 SV Transfer to Complex 39, Pad A May 22, 1969 MSSTransfer to Pad A June 6, 1969 SV Flight Readiness Test (FRT) Completed June 25, 1969 RP-I Loading Completed July 2, 1969 CDDT(Wet) Completed July 3, 1969 CDDT(Dry) Completed July I0, 1969 SV Launch Countdown Started July 16, 1969 SV Launch on Schedule

3-2 the RP-I level display was not a critical measurement, the disposition of the electronics unit was "use as is" However, the level indication was stable during the final 8 hours of countdown.

The RP-I system vent trap closed prematurely during replenish operations at -13 hours, causing entrapped air to be pumped through the S-IC fuel tank. There were no serious consequences. The air, which is filtered to about 50 microns, was immediately vented from the stage. All subsequent system functions were normal, and replenishment was completed sati sfa ctori ly. 3.4.2 LOX Loading

The LOXsystem successfully supported the launch countdown. A premature closure of the S-II stage LOX tank vents during slow fill to 99 percent flight mass caused the LOX loading system to revert at about -6 hours 43 minutes. Recovery procedures were initiated, and flow was reestab- lished at about -6 hours 35 minutes. Launch vehicle loading and replenish operations were completed without further incident. A procedure change will be made to prevent cycling of the tank vents prior to reaching the 99 percent value during future cryogenic loadings.

3.4.3 LH2 Loading The LH2 system supported the launch countdown satisfactorily. A communications problem in the Radio Frequency-Operational Intercom System (RF- OIS) caused a delay in the start of S-II LH2 loading of 25 minutes. The RF-OIS/Pad A fault summary light illuminated at the Launch Control Center (LCC) during LOX loading. This condition could indicate, as a worse case, that pad OIS had switched to batteries or less critical, that an OIS amplifier had switched to secondary. Upon pad entry, an amplifier was found to have automatically switched to secondary; it was reset manually in the PTCRand the fault summary light in the LCCwent off.

During LH2 replenish operations at about -3 hours 20 minutes, a leak de- veloped in the S-IVB stage replenish valve located on LUT level 200. The LH2 system was drained and purged, and the valve bonnet and packing gland bolts were retorqued. No further leakage was detected when LH2 loading operations were resumed at about -2 hours. However, to prevent problem recurrence that could cause countdown delay, the replenish valve was closed and subsequent S-IVB replenishment accomplished manually using the main fill valve in the reduced position.

About 7 minutes after liftoff, during automatic line drain and purge operations, the S-IC liftoff indication was lost causing an LH2 system revert. Drain and purge operations were completed manually using the S-II/S-IVB fill line purge valve. Although this is not the normal manual configuration, a satisfactory purge was obtained. A change in the propellant system logic is presently being considered which will isolate the system from external influence onc_ the liftoff signal is received.

3-3 3.4.4 Auxiliary Propulsion System Propellant Loading

Propellant loading of the S-IVB Auxiliary Propulsion System (APS) was accomplished satisfactorily. Total propellant mass in both modules at liftoff was 184.3 kilograms (406 Ibm) of Nitrogen Tetroxide (N204) and 114.4 kilograms (252 Ibm) of Monomethyl Hydrazine (MMH). 3.5 S-II INSULATION, PURGEANDLEAK DETECTION

The performance of the S-II stage insulation was highly satisfactory. Detailed inspection of all external insulation was conducted by operational television during the countdown and no significant leakage was detected. The total heat leakage through the insulation to the LH2 was within specification limits. Satisfactory pressures and flows were maintained in all purge circuits during countdown. The leak detection system performed satisfactorily throughout the final countdown and contaminant gas concentrations remained within acceptable limits at all times. There were no problems during countdown with the leak detecting selector solenoid valve which presented a minor problem during Countdown Demonstration Test (CDDT).

3.6 GROUND SUPPORT EQUIPMENT (GSE)

3.6.1 Ground/Vehicle Interface

Detailed discussion of the GSE will be contained in the Kennedy Space Center Apollo/Saturn V (AS-506) Ground Systems Evaluation Reports. The performance of all ground systems was highly satisfactory. Overall damage to the pad, LUT and support equipment from the blast and flame impingement was minor. The Holddown Arms (HDA), Tail Service Masts (TSM) and Service Arms (SA) performed within design limits during the launch sequence.

The HDA's were released pneumatically at 0.3 second. HDA No. 1 protective hood did not close and the adjustable head and upper link received some blast damage. However, damage to the interior of HDA No. 1 was not greater than to any other arm. Warpage of the HDA protective hoods was negligible. As on AS-504 launch, the secondary Service Arm Control Switch (SACS) actuator arm on HDA No. 2 was broken off.

TSM retractions were normal and all protective hoods closed properly. The RP-I mast cutoff valve in TSM I-2 opened at liftoff, indicating a loss of valve GN2 control pressure. The cause of pressure loss is being investigated.

SA systems total retract times to safe angle were within specifications. Damage to SA systems was slight. Control console door latches were bent or broken on all SA levels of the LUT; however, provisions incorporated for AS-506 res.trained the doors and prevented their blowing open as had

a-4 occurred on previous launches. Hydraulic oil leakage from SA No. 2 upper and lower hinge areas was detected during postlaunch inspection and was observed to have leaked into SA No. I. Investigation will be conducted to determine the source.

None of the ground/vehicle related problems experienced during launch preparations had sufficient impact such as to constrain the countdown operations. All system repairs and remedial actions were accomplished in parallel with countdown operations. At -13 hours 30 minutes, about 07:31:00 EDT on July 15, 1969, it was discovered that the LCC Data Trans- mission System (DTS) would not synchronize with the DTS transmitter at Pad A. Further investigation revealed severe attenuation of transmitted data. The basic problem was traced to a discrepant patch in the wide- band video distributor. Satisfactory operation was restored at about 19:00:00 EDT of the same day.

3.6.2 MSFC Furnished Ground Support Equipment

Performance of the mechanical and electrical equipment supporting the launch operations was satisfactory. Blast damage to the equipment was considered normal. Minor GSE deviations encountered were as follows:

a, SA No. 1 (S-IC Intertank) umbilical carrier withdrawal time was approximately 0.06 second greater than the specification maximum of 5 seconds. Withdrawal time for this carrier under non-cryogenic conditions, based on the average of results obtained during system revalidation testing, is approximately 3 seconds. Total SA No. 1 retract time to safe angle was 9.9 seconds, which is within the specification limit of 10.5 seconds and was about 3.9 seconds before SA No. 2 retract command. (Failure to achieve SA No. 1 safe angle prior to time for SA No. 2 retract at -16.2 seconds would cause cutoff.) Cause of the slow withdrawal has not yet been determined. Slower than specification withdrawal times were also experienced during the AS-503 and AS-505 launch countdowns. The withdrawal time for the AS-504 launch, although within specification limits, was slower than the average obtained during validation testing under non-cryogenic conditions. Investigation is continuing. b , The GH2 dome regulator in the S-II stage pneumatic servicing console indicated erratic leakage during the -9 hour countdown hold and was replaced with a spare regulator. The new regulator was not adjusted to the high side tolerance of the 810 +.10.3 N/cm z (1175 +_15 psia) setting, as planned. During S-II start tank pressurization, the low regulator setting resulted in the start tank pressures being lower than the desired prelaunch values. At the prelaunch commit point (-33 seconds), S-II Engine No. 1 start tank pressure was 2.8 N/cm 2 (4 psi) below the redline requirement. The countdown was continued since the Central Instrumentation Facility (CIF) observer verified that the measurement was not below redline at -45 seconds.

3-5 The regulator pressures will be set to 827 +_10.3 N/cm2 (1200 +15 psia) for subsequent vehicles and this will alleviate the prelaunch low pressure conditions.

C, The S-II LH 2 heat exchanger delta pressure controller mode of control did not operate properly and the point sensor mode of control was initiated after the beginning of start tank chilldown. This mode of operation was utilized throughout the remaining portion of the countdown. Also, the heat exchanger would not refill properly during the start tank and thrust chamber chilldown sequences. How- ever, the liquid level was sufficient to perform the required stage systems chilldown. The deviation will be investigated.

3.6.3 Camera Coverage

A total of 201 cameras were installed for the AS-506 launch of which 119 were committed to engineering data, and 82 to documentary coverage. Three cameras failed to acquire data. Upon review of film coverage of the GSE at launch, the following conditions were observed:

a. S-II stage forward SA umbilical covers did not secure upon SA withdrawal from the vehicle.

b, HDA No. 1 protective hood failed to close and the other three HDA hoods appeared to close late.

3-6 SECTION4

TRAJECTORY

4.1 SUMMARY

The trajectory parameters from launch to Translunar Injection (TLI) were close to nominal. The vehicle was launched on an azimuth 90 degrees east of north. A roll maneuver was initiated at 13.2 seconds that placed the vehicle on a flight azimuth of 72.058 degrees east of north.

The space-fixed velocity at S-IC Outboard Engine Cutoff (OECO)was 8.5 m/s (27.9 ft/s) greater than nominal. The space-fixed velocity at S-II Out- board Engine Cutoff was 22.8 m/s (74.8 ft/s) less than nominal. The space-fixed velocity at S-IVB first guidance cutoff was 0.2 m/s (0.6 ft/s) less than nominal. The altitude at S-IVB first guidance cutoff was 0.2 kilometer (0.I n mi) lower than nominal and the surface range was 1.7 kilometers (I.0 n mi) less than nominal.

The space-fixed velocity at parking orbit insertion was equal to nominal and the flight path angle was 0.013 degree greater than nominal. The eccentricity was 0.00001 less than nominal. The apogee and perigee were 0.5 kilometer (0.3 n mi) and 0.6 kilometer (0.3 n mi) less than nominal, respectively.

The parameters at translunar injection were also close to nominal. The eccentricity was 0.00029 greater than nominal, the inclination was 0.004 degree greater than nominal, the node was 0.019 degree lower than nominal, and C3 was 16,877 m2/s2 (181,663 ft2/s2) greater than nominal. The space- fixed velocity was 3.2 m/s (10.5 ft/s) greater than nominal and the altitude was 3.1 kilometers (1.6 n mi) less than nominal.

Following Lunar Module (LM) extraction, the vehicle maneuvered to a slingshot attitude frozen rel_tive to local horizontal. The retrograde velocity to achieve S-IVB/IU lunar slingshot was accomplished by a LOX dump, Auxiliary Propulsion System (APS) burn, and LH2 vent The S-IVB/IU closest approach of 3379 kilometers (1825 n mi) above the iunar surface occurred at 78.7 hours into the mission.

The actual impact locations for the spent S-ICi and S-II stages were determined by a theoretical free-flight simulation. The surface range for the S-IC impact point was 0.2 kilometer (0.I n mi) greater than nominal. The surface range for the S-II impact point was 91.7 kilometers (49.5 n mi) less than nominal.

4-I The event times reported in this section reflect the event as seen at the vehicle in order to enable direct comparison with times in the Guidance and Navigation section. 4.2 TRACKINGDATAUTILIZATION

4.2.1 Tracking During the Ascent Phase of Flight

Tracking data were obtained during the period from the time of first motion through parking orbit insertion.

The best estimate trajectory was established by using telemetered guidance velocities as generating parameters to fit data From five different C-Band tracking stations. Approximately 30 percent of the various tracking data was eliminated due to inconsistencies. A comparison of the reconstructed ascent trajectory with the remaining tracking data yielded good agreement. The launch phase portion of the trajectory (liftoff to approximately 20 seconds) was established by constraining integrated telemetered guidance accelerometer data to the early phase of the best estimate trajectory.

4.2.2 Tracking During the Parking Orbit Phase of Flight

Orbital tracking was conducted by the NASAManned Space Flight Network (MSFN). Eight C-Band radar stations furnished data for use in determining the parking orbit trajectory. There were also considerable S-Band tracking data available which were not used due to the abundance of C-Band radar data.

The parking orbit trajectory was obtained by integrating corrected insertion conditions forward to the S-IVB second burn restart preparation event. The insertion conditions, as determined by the Orbital Correction Program (OCP), were obtained by a differential correction procedure which adjusted the estimated insertion conditions to fit the C-Band radar tracking data in accordance with the weights assigned to the data. After all available C-Band radar tracking data were analyzed, the stations and passes providing the better quality data were used in the determination of the insertion conditions.

4.2.3 Tracking During the Injection Phase of Flight

C-Band radar data were obtained from the ship Redstone during the early portion of the injection phase of flight. These tracking data were found to be invalid and were not used in the trajectory determination.

The injection trajectory was established by integrating the telemetered guidance velocity data forward from the restart vector at 9715 seconds (obtained from the parking orbit trajectory) and constraining the end point to the TLI vector at 10,213.03 seconds (obtained from the post TLI trajectory).

4-2 4.2.4 Tracking During the Post Injection Phase of Flight Tracking data from seven C-Band radar stations furnished data for use in determining the post TLI trajectory. The available S-Band tracking data were not used due to the availability of the C-Band radar data.

The post TLI trajectory was obtained by integrating corrected injection conditions forward to S-IVB/Commandand Service Module (CSM) separation. The corrected injection conditions were determined by the same method outlined in paragraph 4.2.2. 4.3 TRAJECTORYEVALUATION

4.3.1 Ascent Trajectory

The vehicle was launched on an azimuth 90 degrees east of north. A roll maneuver was initiated at 13.2 seconds that placed the vehicle on a flight azimuth of 72.058 degrees east of north.

Actual and nominal altitude, surface range, and cross range for the ascent phase are presented in Figure 4-I. Actual and nominal space-fixed velocity and flight path angle during ascent are shown in Figure 4-2.

3000- 240- 1 _ACTUAL ------NOMINAL S-IC OECO 2500- 200- 1 s-if OECO / S-IVB FIRST ECO

E 2000- 160-

ALTITUDE_ ._'_ /

== 1500- _120- p- )" /

cJ -_ 1000- 80- / SURFACE //RANGE_ / //_

500- 40- / _ "__4 -_- CROSS RANGE I i

0_ 0= I 0 lO0 200 300 400 500 600 700 8OO

RANGE TIME, SECONDS

Figure 4-I. Ascent Trajectory Position Comparison

4-3 Comparisons of total inertial accelerations are shown in Figure 4-3. The maximumacceleration during S-IC burn was 3.94 g.

Mach number and dynamic pressure are shown in Figure 4-4. These parameters were calculated using meteorological data measured to an altitude of 56.0 kilometers (30.2 n mi). Above this altitude the measured data were merged into the U.S. Standard Reference Atmosphere.

32- 9000

_ACTUAL 28- 800( ------NOMINAL _7 S-IC OECO J _7 S-II OECO _7 S-IVB FIRST ECO j_ 24- 7000 _ ii ...... / _J

20- '_ 6O

_ SPACE-FIXED I,.- VELOCITY _16- 500 Z

-i- \ ...... / ......

,,12" x 40

I'--- "-r- I

,,,_J LI_. 8- ,t / FLIGHT PATH /...... :______01 I00 / I

-4- I 0 I00 200 300 400 500 600 700 8OO

RANGE TIME, SECONDS

Figure 4-2. Ascent Trajectory Space-Fixed Velocity and Flight Path Angle Comparisons

4-4 4_ Lll ACTUAL ------NOMINAL S-IC OECO ...... H _ s-iI CECO ...... _ s-iI OECO _7 S-IVB FIRST ECO _/ S_7 EMR SHIFT

0 0 _- 2 t,l r_ ..J w I,L ._J

(=3 o

II 8OO 0 I00 200 300 400 50O 600 7O0 RANGE TIME, SECONDS

Figure 4-3. Ascent Trajectory Acceleration Comparison

Actual and nominal values of parameters at significant trajectory event times, cutoff events, and separation events are shown in Tables 4-I, 4-2, and 4-3, respectively.

The free-flight trajectories of the spent S-IC and S-II stages were simulated using initial conditions from the final postflight trajectory. The simulation was based upon the separation impulses for both stages and nominal tumbling drag coefficients. No tracking data were available for verification. Table 4-I presents a comparison of free-flight parameters to nominal at apex for the S-IC and S-II stages. Table 4-4 presents a comparison of free-flight parameters to nominal at impact for the S-IC and S-II stages.

4-5 4.0" 16

A 3.5- 14 --_-- NOMINAL _'_ _ ACTUAL

3.0- 12 DYNAMIC PRESSURE-_

E i (_} z 2.5"

L_ cz') :::D , (.z) ,, 2.0- Z i r-, "-r Y f J _: 1.5- j/ 1.0- / - MACH NUMBER -_ j,_

0.5- / ,-

0 20 40 60 80 I00 120 140 160

RANGE TIME, SECONDS

Figure 4-4. Dynamic Pressure and Mach Number Comparisons

4.3.2 Parking Orbit Trajectory

A family of values for the insertion parameters was obtained depending upon the combination of data used and the weights applied to the data. The solutions that were considered reasonable had a spread of about +250 meters (+820 ft) in position components and +0.7 m/s (+2.3 ft/s) in velocity components. The actual and nominal parking orbit insertion parameters are presented in Table 4-5. The ground track from insertion to S-IVB/CSM separation is given in Figure 4-5.

4.3.3 Injection Trajectory

Comparisons between the actual and nominal space-fixed velocity and flight path angle are shown in Figure 4-6. The actual and nominal total inertial acceleration comparisons are presented in Figure 4-7. Throughout the S-IVB second burn phase of flight, the space-fixed velocity and the flight

4-6 Table 4-I. Comparison of Significant Trajectory Events

EVENT PARAMETER ACTUAL NOMINAL ACT-NOM

First Motion Range Time, sec 0.3 0.3 0.0

Total Inertial Acceleration, m/s 2 10.47 10.61 -0.14 (ft/s 2 ) (34.35) (34.81) (-0.46) (g) (1.07) (i.o8) (-o.oi)

0.7 Mach 1 Range Time, sec 66.3 65.6

Altitude, km 7.8 7.6 0.2 (n mi) (4.2) (4.1) (O.l)

81.3 1.7 Maximum Dynamic Pressure Range Time, sec 83.0

Dynamic Pressure, N/cm 2 3.52 3.47 0.05 (Ibf/ft 2 ) (735.2) (724.7) (10.5)

0.7 Altitude, km 13.6 12.9 (n mi) (7.3) (7.0) (0.3)

Maximum Total Inertial 160.3 1.4 Acceleration: S-IC Range Time, sec 161.7

Acceleration, m/s2 38.61 38.13 0.48 (ft/s z ) (126.67) (125.I0) (1 .57) (g) (3.94) (3.89) (0.05)

460.70 460.26 0.44 S-ll Range Time, sec

Acceleration, m/s 2 17.84 17.99 -0.15 (ft/s 2 ) (58.53) (59.02) (-0.49) (g) (1.82) (l .83) (-o.oi)

699.57 -0.16 S-IVB 1st Burn Range Time, sec 699.41

Acceleration, m/s 2 6.73 6.66 0.07 (ft/s 2 ) (22.08) (2l .85 (0.23) (9) (0.69) (0.68) (o.oi)

I0,204.14 -I.03 S-IVB 2nd Burn Range Time, sec 10,203.11

Acceleration, m/s 2 14.23 14.17 0.06 (ft/s 2 ) (46.69) (46.49) (0.20) (O.Ol) (g) (1.45) (I .44)

Maximum Earth-Fixed 162.3 161.6 0.7 Velocity: S-IC Range Time, sec 5.7 Velocity, m/s 2,402.7 2 _397.0 (18.7) (ft/s) (7,882.9) (7,864.2)

552.52 -3.52 S-ll Range Time, sec 549.00

6,538.8 -23.1 Velocity, m/s 6,515.7 (-75.8) (ft/s) (21,377.0) (21 ,452.8)

709.33 709.49 -o.16 S-IVB Ist Burn Range Time, sec -O.l Velocity, m/s 7,389.5 7,389.6 (24,244.1) (-0.3) (ft/s) (24,243.8)

I0,204.46 -0.96 S-IVB 2nd Burn Range Time, sec 10,203.50

I0,430.2 3.0 Velocity, m/s 10,433.2 (9.9) (ft/s) (34,229.7) (34,219.8

269.1 270.4 -I .3 Apex: S-IC Stage Range Time, sec 115.0 117.3 -2.3 Altitude, km (63.3) (-I .2) (n mi) (62.1

327.4 326.9 0.5 Surface Range, km (176.5) (O.3) (n mi) (176.8)

587.0 593.7 -6.7 S-II Stage Range Time, sec 188.8 189.7 -0.9 Altitude, km (-0.5) (n mi) (lOl .9) (102.4)

1,906.6 -43.7 Surface Range, km 1 ,862.9 (n mY) (1,005.9) (I,029.5) (-23.6)

4-7 Table 4-2. Comparison of Cutoff Events

I PARAMETER ACTUAL NOt41 NALIAC [ NOM ACTUAl NOMINAL ACT-NOM

S-IC CECO (ENGINE SOLENOID) S-IC OECO (ENGINE SOLENOID) Range Time, sec 135.2 135.3 -0.1 ]61.6 161 .I 0.5

Altitude, km 44.0 44.( -0.6 6 6.1 66.7 -0.6 (n mi) (23.8 (24.1) (-0.3) (35.7) (36.0 (-0.3) Surface Range, km 46.4 46.3 0.I 93.6 92.2 1 .4 (n mi) (25.1) (25.0) (0.1) (50.5) (49.8 (0.7) Space-Fixed Velocity, m/s 1 ,979.0 1 ,989.8 -10.8 2,764. 2,765.6 8.5 (ft/s (6,492.8) (6,528.2) (-35.4) (9,068.6) (9,040.7) (27.9)

22.957 Flight Path Angle, deg 23.406 -0.449 19.114 19.635 -0.521

76.315 Heading Angle, deg 76.132 0.183 75.439 75.269i 0.170

Cross Range, km 0.2 0.0 0.2 0.5 0.0 0.5 (n mi) (0.1) (0.0) (0.1)i (0.3 (0.0) (O.3)

Cross Range Velocity, m/s 5.4 -0.2 5.6 12.6 4.3 8.3 (it/s) (17.7) (-0.7) (18.4) (41 .3), (14.1) (27.2)

S-II CECO (ENGINE SOLENOID) S-II OECO (ENGINE SOLENOID)

Range Time, sec 460.6 460.1 O.5 548.2 551.7 -3.5

Altitude, km 180.2 181 . -0.9 187.3 188.0 -0.7 (n mi) (97.3) (97.8 (-0.5 (101.1) (101.5) (-0.4)

Surface Range, km 1 ,114.3, 1,112.5 1 .8 1,617.0 1,640.8 -23.8 (n mi) (601 .7) (600.7 (1.0 (873.1) (886.o (-12.9)

Space-Fixed Velocity, m/s 5,707.5i 5,724.0 -16.5 6,91_.I 6,938.9 -22.8 (it/s) (18,725.4) (18,779.5 (-54.1 (22,69h.6) (22,765.4 (-74.8)

Flight Path Angle, deg 0.897 0.772 0.125 0.608 o.661 -0.053

Heading Angle, deg 79.646 79.658 -0.012 82.389 82.529 -0.140

Cross Range, km 15.0 13.6 1.4 27.4 26.8 0.6 (n mi) (8.1) (7.3) (0.8) (14.8) (14.5) (0.3)

Cross Range Velocity, m/s 111.9 114.5 -2.6 174,1 176.9 -2.8 (ft/s) (367.1) (375.7) (-8.6) (571.2) (58o.4) (-9.2)

S-IVB IST GUIDANCE CUTOFF SIGNAL S-IVB 2ND GUIDANCE CUTOFF SIGNAL

Range Time, sec 699.3 699.5 i -0.2 10,203.0 10,204.1 -].I

Altitude, km 191 .I 191.3 -0.2 320.9 323.8 -2.9 (n mi) (103.2) (103.3) (-0.]) (173.3 (174.8) (-I .5)

Surface Range, km 2,634.0 2,635.7 -1.7 (n mi) (1,422.2) (1,423.2) (-1.O)

Space-Fixed Velocity, m/s 7,791.2 7,791.4 -0.2 10,841.0 10,838.7 2.3 (it/s) (25,561.7) 25,562.3) (-0.6) (35,567.6 (35,560.0) (7.6)

Flight Path Angle, deg 0.015 -0.002 0.017 6.913 6.959 -0.046

Heading Angle, deg 88.416 88.419 -0.003 59.934 59.945 -0.011

Cross Range, km 60.9 59.8 I.I (n mi) (32.9) (32.3) (0.6)

Cross Range Velocity, m/s 274.3 273.3 1.0 (it/s) (899.9) (896.7) (3.2)

Eccentricity 0.975371 0.97542 -0.00005

C3" , m2/s 2 -1,487,528 -1,484,138 -3,390 (ft2/s 2 ) (-16,011,618) (-15,975,128) (-36,490)

Inclination, deg 31.386 31.381 0.005

Descending Node, deg 121.850 121.867 -0.017

* C 3 is twice the specific energy of orbit

C3 = V 2 . _L R where V = Inertial Velocity IJ :: Gravitatinnal Constant R :, Radius vector from center of earth

4-8 Table 4-3. Comparison of Separation Events

PARAMETER ACTUAL NOM I NAL ACT-NOM

S-IC'S-II SEPARATION

Range Time, sec 162.3 161.8 0.5

Altitude, km 66.7 67.4 -0.7 (n mi) (36.0) (36.4) (-0.4)

Surface Range, km 95.1 93.7 1.4 (n mi) (51.3) (5O.6) (0.7)

Space-Fixed Velocity, m/s 2,773.9 2,765.4 8.5 (ft/s) (9,100.7) (9,072.8) (27.9)

Flight Path Angle, deg 19.020 19.533 -0.513

Heading Angle, deg 75.436 75.266 0.170

Cross Range, km 0.5 0.0 0.5 (n mi) (O.3) (0.0) (0.3)

Cross Range Velocity, m/s 12.8 4.4 8.4 (ft/s) (42.0) (14.4) (27.6)

Geodetic Latitude, deg N 28.865 28.865 0.000

Longitude, deg E -79.676 -79.691 0.015

S-II/S-IVB SEPARATION

Range Time, sec 549.0 552.4 -3.4

Altitude, km 187.4 188.1 -0.7 (n mi) (101 .2) (I01.6) (-0.4)

Surface Range, km 1 ,623.4 1 ,645.9 -2_.5 (n mi) (876.6) (888.7) (-12.1)

Space-Fixed Velocity, m/s 6,918.8 6,941 .9 -23.1 (ft/s) (22,699.5) (22,775.3) (-75.8)

Flight Path Angle, deg 0.611 0.653 -0.042

-0.184 Heading Angle, deg 82.426 82.610

0.5 Cross Range, km 27.5 27.0 (n mi) (14.8) (14.6) (O.2)

174.7 -2.6 Cross Range Velocity, m/s 177.3 (ft/s) (573.2) (581.7) (-8.5)

Geodetic Latitude, deg N 31.883 31.921 -0.O38

Longitude, deg E -64.147 -63.913 -0.234

S-IVB/CSM SEPARATION

19 Range Time, sec 11,723 11,704

Altitude, km 7,065.7 6,963.2 102.5 (n mi) (3,815.2) (3,759.8) (55.4)

Space-Fixed Velocity, m/s 7,608.6 7,637.6 -29.0 (ft/s) " (24,962.6) (25,057.7) (-95.1)

Flight Path Angle, deg 45.148 44.922 0.226

Heading Angle, deg 93.758 93.449 0.309

-0.029 Geodetic Latitude, deg N 31.246 31.275

0.483 Longitude, deg E -90.622 -91.105

4-9 Table 4-4. Stage Impact Location

i PARAMETER ACTUAL NOMINAL ACT-NOM

S-IC STAGE IMPACT

Range Time, sec 543.7 546.1 -2.4

Surface Range, km 661 .4 661 .2 0.2 (n mi) (357.1) (357.0) (O.l)

Cross Range, km 8.8 6.3 2.5 (n mi) (4.8) (3.4) (l .4)

Geodetic Latitude, deg N 30.212 30.232 -0.020

Longitude, deg E -74.038 -74.047 0.009

S-II STAGE IMPACT

Range Time, sec 1 ,213.7 1,226.8 -13.1

Surface Range, km 4392.5 4484.2 -91 .7 (n mi) (2371.8) (2421.3) (-49.5)

Cross Range, km 143.0 147.0 (n mi) (77.2) (79.4)

Geodetic Latitude, deg N 31 .535 31.403 0.132

-33.892 -0.952 Longitude, deg E -34.844 path angle were close to nominal with deviations more noticeable towards the end of the time period.

The trajectory and targeting parameters at S-IVB second guidance cutoff and TLI are presented in Tables 4-2 and 4-6, respectively.

4.3.4 Post TLI Trajectory

A family of values for the injection parameters was obtained depending on the combination of data used and the weights applied to the data. The solutions that were considered reasonable had a spread of about ±500 meters (±1640 ft) in position components and ±I.0 m/s (±3.3 ft/s) in velocity components. A comparison of the actual and nominal S-IVB/CSM separation conditions is presented in Table 4-3.

4-10 Table 4-5. Parking Orbit Insertion Conditions

PARAMETER ACTUAL NOMINAL ACT-NOM

Range Time, sec 7O9.3 709.5 -0.2

Altitude, km 191 .I 191 .3 -0.2 (n mi) (103.2) (103.3) (-O.l)

Space-Fixed Velocity, m/s 7793.1 7793.1 0.0 (ft/s) (25,567.9) (25,567.9) (0.0)

Flight Path Angle, deg 0.012 -0.001 0.013

Heading Angle, deg 88.848 88.854 -0.006

Inclination, deg 32.521 32.531 -0.010

Descending Node, deg 123.088 123.100 -0.012

Eccentri city 0.00021 0.00022 -0.00001

Apogee*, km 186.0 186.5 -0.5 (n mi) (100.4) (100.7) (-0.3.)

Perigee*, km 183.2 183.8 -0.6 (n mi) (98.9) (99.2) (-O.3)

Period, min 88.18 88.20 -0.02

Geodetic Latitude, deg N 32.672 32.683 -0.011

Longitude, deg E -52.694 -52.671 -0.023

* Based(3443.934 OnnamSP.herii_ cal earth of radius 6378.165 km

4.3.5 S-IVB/IU Post Separation Trajectory

After final LM separation, the S-IVB/IU was placed on a lunar slingshot trajectory. This trajectory was accomplished by slowing down the S-IVB/IU to make it pass by the trailing edge of the moon and obtain sufficient energy to continue to a solar orbit. This was accomplished by a combination of 108-second LOX dump, 280-second APS burn, and LH 2 vent. A time history of the velocity increase along the vehicle longitudinal axis for the slingshot maneuver is presented in Figure 4-8. Table 4-7 presents a comparison of the actual and nominal velocity increase due to the various

4-11 C) FIRST REVOLUTION

(_) SECOND REVOLUTION

L 2C ,-_ / REIGNITION I/'"I"_

4_ I ,..=I

0 20 40 60 80 I00 120 140 160 180 160 140 120 I00 80 60 40 20

LONGITUDE, deg

Figure 4-5. Ground Track 16- 1

14- _ACTUAL ------NOMINAL S-IVB ENG NE IGNITION v (STDV OPEN) 12- TRANSLUNAR INJECTION / I0- A

,2 b.- SPACE-FIXED

..J VELOCITY ! CD Z / / I 6-- ×

t I

! ! .--J / c_J / / / 2- / FLIGHT _ATH J ANGLE Z J 0= / /

0 I00 200 300 400 500 600 700 800 900 I000

TIME FROM T6, SECONDS

I I I I I

02:35:00 02:38:00 02:41:00 02:44:00 02:47:00 02:50:00

RANGE TIME, HOURS:MINUTES:SECONDS

Figure 4-6. Injection Phase Space-Fixed Velocity and Flight Path Angle Comparisons 1 1.6-

1.4 ACTUAL ------NOMINAL

_7 S-IVB ENGINE IGNITION I .2 _ _7 EMR SHIFT _7 TRANSLUNAR INJECTION

_I.0- E

z 0 Z ! 0 I H <_0.8-- 1-- L_J .J L_ ...1 £.) Lu 4:= £.) I (.) 0.6- 4_

0.4 _

0.2_

O_ I00 200 300 400 500 600 700 800 900 lO00 TIME FROM T6, SECONDS

I IV 02:35:00 02:38:00 02:41:00 02:44:00 02:47:00 02:50:00 RANGE TIME, HOURS:MINUTES:SECONDS

Fi gure 4-7. Injection Phase Acceleration Comparison 0 0 0 0 °. ! .. 0 I ! 0 I r_ I 0 0 I C_J c,'3 0 C) C" Z o. bO

0 .o "--4...,. 0 I_r) °r-- _Z 0 O,J o I I--_" I _0 I

I 0 Z I 0 o. 0 r0 t- I Ii c'J c/3 °t-- r'_ I or) -o Z °. ! 0 z_

I °r-- I I--- {:7) I a 0 Z 0 l O0 o --I I.-- °. l .o w o. r,_ 0 0 % _0 __ 0 0 Z r-- '-r-

0 I,I I-- 0 0 0 .o _.- .0 0 r 0 ..C:

o. I-- 0 0")

0 .ti r'--" 0 c_

CO 0 0 .. 0 I I 0

0 "0 0

or-' LJ-- 0 0 .o "0 0 0 ',,0 o,,I O0 ¢'J O0 i,,--

s/m '39NVHD A±IDOI3A 0 C) z '--" e") • ¢'1 • 0 0 0 (3 0 CO _) sr i I I P- 0 " " " 0 " I--- I ,-- _ 0 0 0 0 _,D " Z (J I v v I I I ¶ p p _ 0 N 0 CO 40 I-- (.- I.-,,. (P v E

u_ r-- I.'1 ,,'_ P'- ,."_ N (_ O_ tO r',, el- ,--, ¢-- O ..J O • N -r-- ,-- 00 r-,, _) _ cO t,O E _J m_" I',. • r-- • _" 0 ¢") CO _0 _" O_ 1---4 ,p- Z _--- - _ (M • t") 4J _: f'l e'l P(M ,--I_ ,-- • r-- Z ("J ¢_(:0 0¢) ¢") r'.- 0 _ ,-- _ CO '- 0 c- 's- 0t- p-- z v o 0 O

c- • ,-- -CO -,_I" "0 O > ....I _) • r-- • i'M P-. • ,r-- ur) t,') (M o Q_ I-- > v v v ,r-) v c'- -J E

4-) c- o c-- 4-) _--. % '-a ch c_ Q; "" E_ c- e_ .i-) ,t- ..q-

S. E u Q_ E 4..) O

I 4-) O_ c- >_ -q- _) O :h _n I-- Iu _'-,-" "0 0') ,r-- u i/I _ E .It r_ u I--- P" -_..lla -_ .ll-! o I-- ,,y c 0 Q C_) I-- P" I_. t=: 0 0 E ,-- " Z >_ A N E L E 0 0"1 -1,... N I/1 i= =1 =_ E ¢1 -.-., Q; 0 E " X ¢1 _ .,- e- {j Ul "-. i =l x _ E c,; I.-- "CI I.i. 4._ :3 I 4; JE .P -,-- (1) E r_ .,J

•_" m O {n U " I.-- O P- e.I E Q; U ¢_) ._1 Ia.. "I" _ O I._ U phases of the maneuver. The major error contribution in total velocity increase is due to the resulting 7.3 m/s (24.0 ft/s) from the Continuous Vent System (CVS) as compared to 3.5 m/s (11.5 ft/s) for the predicted value. Figure 4-9 presents the resultant conditions for various velocity increases at the given attitude of the vehicle for the maneuver.

The S-IVB/IU closest approach of 3379 kilometers (1825 n mi) above the lunar surface occurred at 78.7 hours into the mission. The trajectory parameters were obtained by integrating forward a vector (furnished by Goddard Space Flight Center (GSFC) which was obtained from Unified S-Band (USB) tracking data during the active lifetime of the S-IVB/IU. The actual and nominal conditions at closest approach are presented in Table 4-8. Figure 4-10 illustrates the influence of the moon on the S-IVB/IU energy (velocity) relative to the earth, particularly as the spent stage passes through the lunar sphere of influence. Some of the heliocentric orbit parameters of the S-IVB/IU are presented in Table 4-9. The same parameters for the earths orbit are also presented for comparison.

4-17 ATTITUDE (LOCAL HORIZONTAL REFERENCE SYSTEM)

218 ° PITCH 0 ° YAW 170 ° ROLL

EARTH LUNAR EARTH-MOON EARTH CAPTURE IMPACT ESCAPE CAPTURE

0 5 I0 15 20 25 35

VELOCITY CHANGE, m/s NOMINAL ACTUAL (31 .5) (36.3)

Figure 4-9. Trajectory Conditions Resulting from Slingshot Maneuver Velocity Increments

4TH DAY 2,4"

I 5TH_ DAY

I 16TH DAY 10TH DAY 21 ST DA_ LUNAR I 1 !i I 2o12/il E.co0,TERII I I I 1.6 1 I I I I t l I

_- 30 I 1.2= I _.1 I 48 HR I O.

72 HR

0.4-

0,0 0.0 0.2 0.4 0.6 0.8 l .0 1 .2 1 .4 1 .6 1 .8

DISTANCE (EARTH-VEHICLE), 106 km

Figure 4-10. S-IVB/IU Velocity Relative to Earth Distance

4-18 Table 4-8. Comparison of Lunar Closest Approach Parameters

PARAMETER ACTUAL NQMINAL ACT-NOM

Lunar Radius, km 5117 370O 1417 (n mi) (2763) (1998) (765)

Altitude Above Lunar Surface, km 3379 1962 1417 (n mi) (1825) (I059) (765)

Range Time, hr 78.7 78.4 0.3

Velocity Increase Relative to Earth from Lunar Encounter, km/s 0.680 0.860 •-0.180 (n mi/s) (0.367) (0.464) (-0.097)

Table 4-9. Heliocentric Orbit Parameters

PARAMETER S-IVB/IU EARTH

Semimajor Axis, lO 6 km 143.08 149.00

(lO 6 n mi) (77.26) (80.45)

Aphelion, lO 6 km 151 .86 151 .15

(lO 6 n mi) (82.00) (81 .61)

Perihelion, 106 km 134.30 146.84

(lO 6 n mi) (72.52) (79,29)

Inclination,* deg 0.3836 0.0000

Period, days 342 365

* Measured with respect to the ecliptic.

4-19/4-20

SECTION5

S-IC PROPULSION

5.1 SUMMARY

All S-IC propulsion systems performed satisfactorily and the propulsion performance level was very close to nominal. Stage site thrust (averaged from liftoff to Outboard Engine Cutoff [OECO]) was 0.62 percent lower than predicted. Total propellant consumption rate was 0.40 percent lower than predicted with the total consumed Mixture Ratio (MR) 0.I0 percent lower than predicted. Specific impulse was 0.16 percent lower than predicted. Total propellant consumption from Holddown Arm (HDA) release to OECOwas low by 1.12 percent. Center Engine Cutoff (CECO)was initiated by the Instrument Unit (IU) at 135.20 seconds as planned. OECO,initiated by LOX low level sensors, occurred at 161.63 seconds which was 0.55 second later than predicted. This is a small difference compared to the predicted 3-sigma limits of +_3.74 seconds. The LOX residual at OECOwas 18,041 kilograms (39,772 Ibm) compared to the predicted 18,177 kilograms (40,074 Ibm). The fuel residual at OECOwas 13,954 kilograms (30,763 Ibm) compared to the predicted 14,354 kilograms (31,645 Ibm). 5.2 S-IC IGNITION TRANSIENTPERFORMANCE

The fuel pump inlet preignition pressure was 31.6 N/cm2 (45.9 psia) and within F-I Engine Model Specification limits of 30.0 to 75.8 N/cmL (43.5 to II0 psia). The fuel pump inlet preignition temperatures were not available since these measurements were deleted from the S-IC-6 and subsequent stages. The LOX pump inlet preignition pressure and temperature were 58.5 N/cm2 (84.8 psia) and 96.1°K (-286.7°F) and were within the F-I Engine Model Specification limits as shown in Figure 5-I. Engine startup sequence was nominal. A I-2-2 start was planned and attained. Engine position starting order was 5, I-3, 4-2. Two engines are considered to start together if their combustion chamber pressures reach 68.9 N/cm2 (I00 psig) in a lO0-millisecond time period.

5-I LOX PUMP INLET PRESSURE, psia 60 8O I00 120 140 160 180 110

-265

106 PREDI CTED- --STARTING o REGION -275 1o2 F-- ...... I o I = I! ACTUAL I 96.1°K (-286.7°F I :EZ 58.5 N/cm2 (84.8 psia) i,i I _- 98 I I I -285 I I L I I

9 _ I I

x o [ I -295 __1 i 90":

..... "_ ACCFPTABLE STARTING -- REGION 86 .... -305

4O 5O 60 70 80 90 I00 110 120 130 LOX PUMP INLET PRESSURE, N/cm 2

Figure 5-I. S-IC LOX Start Box Requirements

Figure 5-2 shows the thrust buildup of each engine indicative of the successful I-2-2 start. The shift in thrust buildup near the 5,250,000 Newtons (1,180,000 Ibf) level on the outboard engines is caused by ingestion of helium from the LOX prevalves during startup. The thrust shift is absent on the center engine since the POGO suppression helium accumulator system is not used on this engine. Engine combustion chamber pressure oscillograms show 79- to 80-hertz oscillations of approximately 445,000 Newtons (I00,000 Ibf) peak-to-peak amplitude during buildup. These oscillations are characteristic of normal F-I engine thrust buildup. Engines No. 1 and 5 show normal inertial surge chamber pressure spikes of approximately 48.3 N/cm 2 (70 psi) and 50.3 N/cm2 (73 psi), respectively, at 3.45 seconds after their individual start solenoids were energized. Engine No. 4 data indicate a large chamber pressure spike (approximately 80 percent of the mainstage level) at 3.42 seconds after engine No. 4 start solenoid energization. The unusual magnitude of this spike is believed to have been the result of a data,problem and is a characteristic of the flight pressure transducer. Static firings of the F-I engines have exhibited similar pressure spikes (measured with the flight pressure transducer) during the buildup transient, but failed to indicate the same

5-2 7.0 ...... - ...... _ _": : ..... _ .",<_I• _.- _-'---_ -_'_. -I .50

' i ENGINE NO. 5-_. // /

• 6.0' - _ i ,

,*" _.I i /i "I .25

ENGINE NO. I_-____ _ / 5.0 ." - J : l

. //I -I .00 • _--ENGINE NO. 2 4- z 4.0 .:

-0.75 _3.0 _-- .: / I I • . _--ENGINE NO. 4 I-.-

0.50 : / 2.0 " II

I I :" : I

1.0 i i I_ ENGINE NO., 3-_ _ _I;i 0.25 1 " 0 ,-...... - __=___s.,,_...._ 0 -4.5 -4.0 -3.5 -3.0 -2.5 -2.0 -I .5 -I .0 -0.5 0 RANGE TIME, SECONDS

Figure 5-2. S-IC Engines Buildup Transients

spike on high frequency type ground firing instrumentation. The pressure spike has, therefore, been omitted from the thrust buildup curve shown in Figure 5-2.

The best estimate of propellants consumed between ignition and HDA release was 39,374 kilograms (86,803 Ibm). The predicted consumption was 38,913 kilograms (85,790 Ibm). Propellant loads at HDA release were 1,468,594 kilograms (3,237,697 Ibm) for LOX and 637,830 kilograms (1,406,175 Ibm) for fuel.

5.3 S-IC MAINSTAGE PERFORMANCE

S-IC stage propulsion performance was satisfactory. Site performance was very close to the predicted level as can be seen in Figure 5-3. The stage site thrust (averaged from liftoff to OECO) was 0.62 percent lower than predicted with the total propellant consumption rate 0.40 percent lower than predicted and the total consumed propellant MR 0.I0 percent lower than predicted and the specific impulse 0.16 percent lower than predicted. Total propellant consumption from HDA release to OECO was low by 1.12 percent.

The F-I engines performance levels during the AS-506 flight showed the smallest deviations from predicted levels of any S-IC flight.

5-3 • i7-_

LOX/FUEL MIXTURE RATIO STAGE THRUST, 106 N N N N r_ r_ N E l

_m m_

o_ rm_

,..J° \ c- 8 \ cD 8 , _ \

I I,-,q 2_ | z

•_ co co o o t. _ L.

e- STAGE THRUST, 106 Ibf

O0 STAGE SPECIFIC IMPULSE, 103 N-s/kg STAGE FLOWRATE, 104 kg/s ..J, o N ¢m _ ?0

--h o \

0.; \

I \ ,t (/) ° \ \

"I N e,_ w Po N r-_ N ,m, o_ CO

STAGE SPECIFIC IMPULSE, 102 Ibf-s/Ibm STAGE FLOWRATE, 10 4 Ibm/s For comparing F-I engine flight performance with predicted performance, the flight performance has been analytically reduced to standard conditions and compared to the predicted performance which is based on ground firings and also reduced to standard conditions. These values are shown in Table 5-I at the 35- to 38-second time slice Individual engine deviations from predicted thrust ranged from 0,662 percent lower tengine No. 5) to 0.527 percent higher (engine No. 4). Individual engine deviations from predicted specific impulse ranged from 0.114 percent lower (engine No. 5) to 0.038 percent higher (engines No. 1 and 4).

5.4 S-IC ENGINE SHUTDOWN TRANSIENT PERFORMANCE

CECO was initiated by a signal from the IU at 135.20 seconds as planned. OECO, initiated by LOX low level sensors, occurred at 161.63 seconds which was 0.55 second later than predicted. This is a small difference compared to the predicted 3-sigma limits of +3.74 seconds. Most of the OECO deviation can be attributed to lower than predicted thrust, specific impulse, and propellant loads.

Thrust decay of the F-I engines was nominal.

Table 5-I. S-IC Engine Performance Deviations

AVERAGE RECONSTRUCTION DEVIATI ON DEVI AT I ON PARAI.IETER ENGINE PREDICTED ANALYSIS PERCEHT PERCENT

Thrust 6727 (1512) 6740 (1515) 0.198 103 N (I03 Ibf) 6695 (1505) 6674 (1500) -0.332 6717 (1510) 6725 (1512) 0.132 -O.027 6748 (1517) 6783 (1525) 0.527 6717 (1510) 6674 (1500) -0.662

Specific Impulse 2598 (264.9) 2599 (265.0) 0.038 H-s/kg (Ibf-s/Ibm) 2599 (265.0) 2598 (264.9) -0.038 2596 (264.7) 2596 (264.7) 0 -0.015 2594 (264.5) 2595 (264.6) 0.038 2587 (263.8) 2584 (263.5) -0.114

Total Flowrate 1 2589 (5708) 2594 (5718) 0.175 kg/s (Ibm/s) 2 2576 (5679) 2569 (5664) -O.264 3 2587 (5703) 2590 (5711) 0.140 -0.025 4 2602 (5737) 2613 (5761) 0.418 5 2597 (5725) 2582 (5691) -0.594

Mixture Ratio 2.258 2.255 -0.133 LOX/Fuel 2.244 2.241 -0.134 2.262 2.259 -0.133 -0.133 2.254 2.251 -0.133 2.282 2.279 -0.131

NOTE: Performance levels were reduced to standard sea level and pump inlet conditions at 35 to 38 seconds.

5-5 Engine cutoff impulse was approximately 10,612,096 N-s (2,385,694 Ibf-s) or II percent higher than predicted for the outboard engines and approximately 2,659,605 N-s (597,903 Ibf-s) or 7 percent lower than predicted for the center engine. The impulse values stated for the outboard engines are for the period from cutoff signal to stage separation, and the impulse value for the center engine is for the period from cutoff signal to zero thrust of the center engine. The flight cutoff impulse is based on chamber pressures. At cutoff, chamber pressure was high for engines No. I, 3 and especially 4, and low for engine No. 5. These chamber pressure deviations yielded sufficient thrust to account for the cutoff impulse deviations. 5.5 S-IC STAGEPROPELLANTMANAGEMENT

The S-IC does not have an active Propellant Utilization (PU) system. Minimum residuals are obtained by attempting to load the mixture ratio expected to be consumed by the engines plus the predicted unusable residuals. An analysis of the usable residuals experienced during a flight is a good measure of the performance of the passive PU system.

OECOwas initiated by the LOX low level sensors as planned, and resulted in residual propellants being very close to the predicted values. The residual LOX at OECOwas 18,041 kilograms (39,772 Ibm) compared to the predicted value of 18,177 kilograms (40:074 Ibm). The fuel residual at OECOwas 13,954 kilograms (30,763 Ibm) compared to the predicted value of 14,354 kilograms (31,645 Ibm). A summary of the propellants remaining at major event times is presented in Table 5-2.

Table 5-2. S-IC Stage Propellant Mass History

LEVEL SENSOR EVENT PREDICTED DATA RECONSTRUCTED LOX FUEL LOX FUEL LOX FUEL

Ignition kg 1,500,418 646,854 646,323 1,499,479 646,319 Command (Ibm) (3,307,854) (1,426,070) (1,424,899) (3,305,786) (1,424,889)

Holddown kg 1,469,966 638,393 1,468,792 637,386 1,468,594 637,830 Arm Release (Ibm) (3,240,719) (1,407,415) 3,238,132) (1,405,195) (3,237,697) (1,406,175)

CECO kg 211,g56 97,465 217,230 99,475 216,633 99,059 (Ibm) (467,282) (214,874) (478,911) (219,304) (477,594) (218,389)

OECO kg 18,177 14,354 19,009 14,202 18,041 1 3,954 (Ibm) (40,074) (31,645) (41,908) (31 ,309) (39,772) (30,763)

Separation kg 15,594 13,263 15,651 12,705 (Ibm) (34,377) (29,241) (34,504) (28,008)

Zero Thrust kg 15,406 13,063 15,408 12,517 (Ibm) (33,965) (28,800) (33,970) (27,595)

NOTE: Predicted and reconstructed values do not include pressurization gas so they will compare with level sensor data.

5-6 5.6 S-IC PRESSURIZATIONSYSTEMS

5.6.1 S-IC Fuel Pressurization System

The fuel tank pressurization system performed satisfactorily keeping ullage pressure within the acceptable limits during flight. Helium Flow Control Valves (HFCV's) No. 1 through 4 opened as planned and HFCVNo. 5 was not required.

The low flow prepressurization system was commandedon at -97 seconds. High flow pressurization, accomplished by the onboard pressurization system, performed as expected. HFCVNo. 1 was commandedon at -2.7 seconds and was supplemented by the high flow prepressurization system until umbilical disconnect.

Fuel tank ullage pressure was within the predicted limits throughout flight as shown in Figure 5-4. HFCV's No. 2, 3, and 4 were commanded open during flight by the switch selector within acceptable limits. Helium bottle pressure was 2137 N/cm2 (3100 psia) at -2.8 seconds and decayed to 331N/cm 2 (480 psia) at OECO. Total helium flowrate and heat exchanger performance were as expected.

Fuel pump inlet pressure was maintained above the required minimum Net Positive Suction Pressure (NPSP) during flight.

5.6.2 S-IC LOX Pressurization System

The LOX pressurization system performed satisfactorily and all performance requirements were met. The ground prepressurization system maintained ullage pressure within acceptable limits until launch commit. The onboard pressurization system subsequently maintained ullage pressure within the GOXFlow Control Valve (GFCV) band during flight.

The prepressurization system was initiated at -72 seconds. Ullage pressure increased to the prepressurization switch band and flow was terminated at -57 seconds. The low-flow system was cycled on two additional times at -40 and -17 seconds. At -4.7 seconds the high-flow system was commanded on and maintained ullage pressure within acceptable limits until launch commit.

The LOX tank ullage pressure during flight, shown in Figure 5-5, was maintained within the required limits throughout flight by the GFCV. The maxi- mumGOXflowrate to the tank (at CECO)was 24.9 kg/s (55.0 Ibm/s). The heat exchangers performed as expected.

The LOX pump inlet pressure met the minimum NPSPrequirement throughout flight. The engine No. 5 LOX suction duct pressure decayed after CECO similar to previous flights as shown in Figure 5-6. The cause of these decays is still unknown.

5-7 _7HFCV NO. 2 OPEN XTHFcv N(). I OPEN I 22 '_THFCV NO. 3 OPEN 32 _THFCV NO. 4 OPEN /-'PREDICTED MAXIMUM / I I \ / I s

\ I \ __.I 30 ._ -_. \ / \I "_I

F 28 m

e_ Q..

.._/ ._ \/ v 26 _

z z \ I--- I-- '\ \ 24 _

16 _ \ / \ / \ J _ A / _ ..... _ 22 \_" I I / _'" PREDICTED MINIMUM 14 _ _7, I I T _z 0 20 40 60 80 1O0 120 140 160 RANGE TIME, SECONDS

Figure 5-4. S-IC Fuel Ullage Pressure

5.7 S-IC PNEUMATIC CONTROL PRESSURE SYSTEM

The control pressure system functioned satisfactorily throughout the S-IC flight.

Sphere pressure was 2151N/cm 2 (3120 psiaZ at liftoff and remained steady until CECO when it decreased to 2068 N/cmL (3000 psia). The decrease was due to center engine prevalve actuation. There was a further decrease to 1810 N/cm 2 (2625 psia) after OECO.

The engine prevalves were closed after engine cutoff as required. The engine No. 5 prevalves closed at approximately 137 seconds. The prevalves for the other four engines closed at approximately 163 seconds.

5.8 S-IC PURGE SYSTEMS

Performance of the S-IC purge systems was satisfactory during the flight.

The turbopump LOX seal purge storage sphere pressure was within the limits of 1862 to 2275 N/cm2 (2700 to 3300 psia) until ignition and 2275 to 689 N/cm 2 (3300 to I000 psia) from liftoff to cutoff. The radiation calorimeter purge system was not installed on S-IC-6 nor subsequent vehicles.

5-8 AdlVE #RESSDRE 3O RELIEF SWITCH BAND DURING FLIGHT

...... A-MECHANICAL REI IEF BAND ...... "...... /J 16.2 to 17.6 N/cm2 26 c_ ...... "..... "-. _'""//(23.5"'-* I [ to 25.5 psig) --RELIEF SWITCH BAND ._ E t) z

1,1 r_ 22 tz_ cz_ w c_ ...... -30 o_ (i_

L._ N CD "__.__ _ _-- PREPRESS16.7 to 18.3SWITCHN/cm2BAND -.,-.z_,/_/j ,_,_ >!." =__J "" (24.2 to 26.5 psia) _/7_>-F ...... __. z - I.-, I-- x

.if __i! " _

-AS-5O6 FLIGHT DATA PREDICTED LIMITS I .15 0 20 40 60 80 I00 120 140 160 P,ANGE TIME, SECONDS

Figure 5-5. S-IC LOX Tank Ullage Pressure

5.9 S-IC POGO SUPPRESSION SYSTEM

The POGO Suppression system performed satisfactorily during S-IC flight.

Outboard LOX prevalve temperature measurements indicated that the prevalve cavities were filled with helium prior to liftoff as planned. The measurements in the outboard prevalves went cold momentarily at liftoff indicating LOX sloshed on the probes. They remained warm throughout flight, indicating helium in the prevalves. At cutoff, the increased pressure forced LOX into the prevalves once more. The two measurements in the center engine prevalve indicated cold, which meant LOX was in this valve, as planned.

5-9 I S_7CECO _- AS-505 _70ECO 120

80 _.=....- _, I00 c_J E ,I (.} o _,, _-AS-504 z 60 r_

CZ_ 8O CZ) rY _" " /-AS-506 r_ cz) I CZ_ c_ 0 r_

F-- 0 6O -_ 40 0 r-_ F-- Z 0 0 _=_ CZ) F- (..) X 0 cz_ x C) " 20

_7 _7 0 0 -5 0 5 I0 15 20 25 30 TIME FROM CECO, SECONDS ! ! I I I I I I 130 135 140 145 150 155 160 165 RANGE TIME, SECONDS

Figure 5-6. S-IC LOX Suction Duct Pressure, Engine No. 5

5-10 SECTION 6

S-II PROPULSION

6.1 SUMMARY

The S-II propulsion system performed satisfactorily throughout the flight. As sensed at the engines, Engine Start Command (ESC) occurred at 163.04 seconds and Outboard Engine Cutoff (OECO) at 548.22 seconds with an operation time of 385.18 seconds or 4.0 seconds shorter than predicted. Due to high amplitude low frequency oscillations on the AS-503 and AS-504 flights, the center engine was shut down early as on AS-505 and successfully avoided these oscillations. Center Engine Cutoff (CECO) occurred at 460.62 seconds. Total stage thrust, as determined by computer analysis of telemetered propulsion measurements, at 61 seconds after S-II ESC was 0.20 percent below predicted. Total propellant flowrate (including pressurization flow) was 0.13 percent below predicted and stage specific impulse was 0.07 percent below predicted at this time slice. Stage propellant Mixture Ratio (MR) was 0.36 percent above predicted.

The propellant management system performance was satisfactory. The system was similar to AS-505 in that it also used open-loop control of the engine Propellant Utilization (PU) valves. On AS-506, however, the Instrument Unit (IU) command to shift Engine Mixture Ratio (EMR) from high to low was initiated upon attainment of a preprogramed stage characteristic velocity as sensed by the Launch Vehicle Digital Computer (LVDC). An IU timed command served this function on AS-505. The IU EMR shift command occurred 6 seconds later than predicted and this deviation was due mainly to improper scaling in the LVDC velocity computations. The actual shift from high to low EMR occurred 9.5 seconds late when compared with the final propulsion prediction. The additional 3.5 seconds resultl from a propulsion and characteristic velocity presetting mismatch that was known prior to flight. Future preflight operational trajectory events, IU programed commands, and S-II propulsion prediction events will be reviewed for compatibility.

OECO, initiated by the LOX low level cutoff sensors, was achieved following a planned 1.5-second time delay. A small engine performance decay was noted just prior to cutoff similar to AS-505, but was less severe than that observed on AS-504 due to only four engines operating at cutoff. Residual propellant remaining in the tanks at OECO signal was 3388 kilograms (7471 Ibm) compared to a prediction of 2623 kilograms (5783 Ibm).

6-I The performance of the LOX and LH2 tank pressurization systems was satisfactory. Ullage pressure in both tanks was more than adequate to meet engine inlet Net Positive Suction Pressure (NPSP) requirements throughout mainstage. As commandedby the IU, step pressurization occurred at 261.6 seconds for the LOX tank and 461.6 seconds for the LH2 tank.

The engine servicing system performed satisfactorily except that the engine No. 1 start tank pressure was 2.8 N/cm2 (4 psi) below redline at prelaunch commit (-33 seconds). This low pressure was caused by a lower than planned setting of the Ground Support Equipment (GSE) regulator supplying hydrogen to the start tank. Corrective action being proposed includes increasing the nominal setting of the GSE regulator and relaxing the prelaunch commit redline to more closely approximate actual requirements. All start tank pressures and temperatures were well within requirements at S-II ESC.

The recirculation, pneumatic control and helium injection systems all performed satisfactorily.

6.2 S-II CHILLDOWN AND BUILDUP TRANSIENT PERFORMANCE

The prelaunch servicing operations satisfactorily accomplished the engine conditioning requirements. Thrust chamber temperatures were within predicted limits both at launch and engine start. The thrust chamber temperatures ranged between I01 and II9°K (-278 and -245°F) at prelaunch commit and 131 and 150°K (-223 and -190°F) at engine start. Thrust chamber temperature warmup rate during S-IC boost agreed closely with those experienced on previous flights.

Engine start tank temperatures at the conclusion of chilldown ranged between 95 and IO0°K (-289 and -280°F) and were similar to AS-505. All start tank temperatures and pressures were within the prelaunch and engine start boxes, as shown in Figure 6-I, with the exception that engine No. 1 start tank pressure was 2.8 N/cm2 (4 psi) low at prelaunch commit (-33 seconds).

The low start tank pressures at -33 seconds resulted from the start tanks being pressurized at 783 to 792 N/cm2 (!135 to 1148 psia) instead of the required 810 +_10.3 N/cm 2 (1175 +_15 psia). It had been planned to set both the GSE S-II pneumatic console dome regulator and the start tank supply regulator at the high side of the tolerance. The dome regulator was replaced during the -9 hour launch countdown hold Without adjustment to the high limit (refer to paragraph 3.6.2). Another factor contributing to the low start tank pressures was that the pressure gauge used to set the regulators was reading approximately 7.6 N/cm2 (II psi) high. It is planned to revise the pressurization regulator settings to provide a higher pressure level for subsequent stages. It has also been recommended that the minimum pressure line of the prelaunch redline box be lowered approximately 6.9 N/cm2 (I0 psi). Review of all previous launch data indicates a lower prelaunch pressure is compatible with the engine start box.

6-2 START TANK TEMPERATURE, °F -300 -250 -200 -I 50 1050 , , 0 ENG. NO. 1 500 _7 ENG. NO. 2 [] ENG. NO. 3 ENGINE START BOX. I000-- 0 ENG. NO. 4 - PRELAUNCHIBOX- \ ENG. NO. 5 /- 400 "_ =" 950 I_ ...... S_ °

_ t I mm 900 .,-" o I 300 _- I S-II ESC '

"_ j ..J

_ I--- __ 850 ', o [] _....:_---'_ -

i.--

800 PRELAUNCH (-33 SEC)

750 II00 80 90 I00 II0 120 130 140 150 160 170 180

START TANK TEMPERATURE, °K

Figure 6-I. S-II Engine Start Tank Performance

All engine helium tank pressures were within the prelaunch and engine start limits of 1931 to 2379 N/cm z (2800 to 3450 psia). The helium supply line was manually vented at -277 seconds versus being vented at -30 seconds on previous launches. This allowed adequate time to monitor for leakage prior to the -19 second launch commit. No pressure decay of any significance occurred during this time period.

Engine No. 2 helium tank pressure decayed at a sharper rate than expected after S-II ESC. The decay assumed a more normal rate after approximately 30 seconds of operation. This condition has occurred on previous flights and has been coincident with shifts in the engine helium regulator outlet pressure. Engine regulator outlet pressure measurement was not provided on AS-506 so it can only be assumed that a regulator outlet pressure shift also occurred. On AS-505 flight, engine No. 5 regulator outlet pressure shifted from 281 to 276 N/cm z (408 to 400 psia) at approximately 63 seconds after ESC. On AS-504 flight, engine No. 3 regulator outlet pressure shifted from 279 to 276 N/cm2 (405 to 400 psia) at approximately 43 seconds after ESC. Between ESC and regulator shift the decay rates were higher than expected, but following the shiftthe decay rates of all engines were comparable.

6-3 The higher than expected helium tank decay rates experienced to date are not critical for the S-II mission. Even if the initial decay rate continued throughout S-II burn, the supply pressure would be adequate to meet system demandswith sufficient margin. The cause of this deviation has been assessed as internal leakage through the engine helium regulator.

The LOX and LH2 recirculation systems used to chill the feed ducts, turbo- pumps, and other engine components performed satisfactorily during prelaunch and S-IC boost. Engine pump inlet temperatures and pressures at engine start were well within the requirements as shown in Figure 6-2. The LOX pump discharge temperatures at ESCwere 7.5 to 8.9°K (13.5 to 16.1°F) subcooled, which is well below the 1.7°K (3°F) subcooling requirement.

Prepressurization of the propellant tanks was satisfactorily accomplished. Ullage pressures at S-II ESCwere 26.9 N/cm2 (39 psia) for LOX and 19.6 N/cm2 (28.5 psia) for LH2. S-II ESCwas received at 163.04 seconds and the Start Tank Discharge Valve (STDV) solenoid activation signal occurred 1.0 second later. The engine thrust buildup was satisfactory and was within the required thrust buildup envelope. The stage thrust reached mainstage level at 166.2 seconds. Engine thrust levels were between 861,496 and 895,080 Newtons (193,672 and 201,222 Ibf) prior to "High EMRSelect" commandat 168.5 seconds. 6.3 S-II MAINSTAGEPERFORMANCE

Stage performance during the high EMRportion of the flight was very close to predicted as shown in Figure 6-3. At a time slice of ESC+61 seconds, total vehicle thrust was 5,141,516 Newtons (1,155,859 Ibf) which is only 10,094 Newtons (2269 Ibf) or 0.20 percent below the preflight prediction. Total propellant flowrate (including pressurization flow) was 1239 kg/s (2731 Ibm/s) which was 0.13 percent below prediction. Stage specific impulse,,including the effect of pressurization gas flowrate, was 4150.2 N-s/kg (423.2 Ibf-s/Ibm) which is 0.07 percent below the predicted level. Stage propellant MRwas 0.36 percent above prediction.

At ESC+297.58 seconds (460.62 seconds) the center engine was shut down in order to prevent buildup of the low frequency oscillations that were observed on AS-503 and AS-504. This action reduced total vehicle thrust by 1,031,685 Newtons (231,932 Ibf) to a level of 4,093,107 Newtons (920,167 Ibf). Of this total, a thrust reduction of 1,017,255 Newtons (228,688 Ibf) was directly due to CECOand the remaining 14,430 Newtons (3244 Ibf) decrease resulted from the sum effect of fuel step pressurization (ESC +298.6 seconds) and loss of acceleration head. The shift from high to low EMRoperation occurred at approximately 335 seconds after ESC. The change of EMRresulted in further thrust reduction, and at ESC+351 seconds the total vehicle thrust was 3,082,769 Newtons (693,034 Ibf); thus a decrease in thrust of 1,010,338 Newtons (227,133 Ibf) is indicated between high and the average low EMRoperation.

6-4 :1o '3WNIV_d3dN31 137NI dNNd gH7 :1o '3WfllVd3dN31 137NI dNfld X07

CO 0 _ _ _ CO 0 _ re:l- _0 CO r-- r-- r-- Od C_I C_I CO O0 CO CO O_ 0"_ Oh Oh Oh E

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_,o '3_fllV_3dN3I 137NI dNnd gH7 )Io '_fllY_l_dN31 137NI dNfld X07 X_7 S-ll ESC V S-II EMRSHIFT _7S-II ESC _S-II EMRSHIFT .... PREDICTED .... PREDICTED X_ S-ll CECO _S-II OECO _RECONSTRUCTED _7S-II CECO _7S-II OECO _RECONSTRUCTED 5.4 1400 - 1.16 3OOO 5.0 1300 2800 1.08 4.6 1200 -?. 1.00 2600 z .Q 4.2 II00 0.92 F-- 2400 _

_c 0 I000 3.8 0.84 22OO_ I r_ #- i I F-- 3.4 0.76 0 900 20O0_ I i L,. I 0.68 1800 3.0 8OO L 0.60 "_-'-- 1600 2.6 700

I O. 52 1400 O_ 2.2 6OO

4.4 5.6 ku --J _Z3 o F 440 2 4.3 < 5.2 I

Q.) I 430 Lz- I _- 4.2 4.8 _E __.J I

420 _ -_. ta.I U- I N_' 4.1 _-.. 4.4 _m i L_._ D---O c,3 __1 410 (z3__ 4.0 4.0 0 50 1O0 150 200 250 300 350 400 0 50 1O0 150 200 250 300 350 400 TIME FROMESC, SECONDS TIME FROMESC, SECONDS V , , , i , VV,_ , , , , , W'_' 150 200 250 300 350 400 450 500 550 150 200 250 300 350 400 450 500 550 RANGETIME, SECONDS RANGETIME, SECONDS

Figure 6-3. S-ll Steady State Operation Similar to AS-505 flight, the deviation of actual from predicted performance remained small at the lower mixture ratio levels. At ESC +381 seconds, total thrust was 3,059,402 Newtons (687,781 Ibf) at an EMR of 4.29. Vehicle thrust and propellant flowrate deviations at this time were 18,683 Newtons (4200 Ibf) and 5.1 kg/s (11.2 Ibm/s), respectively.

Individual J-2 engine data, excluding the effects of pressurization flowrate, are presented in Table 6-I for the ESC +61-second time point. Very good correlation between prediction and flight is indicated by the small magnitude of the deviations. Flight data reconstruction procedures were directed toward matching the engine and stage acceptance specific impulse values while maintaining the engine flow and pump speed data as a baseline.

Data presented in Table 6-I are actual flight data and have not been adjusted to standard J-2 engine conditions. Considering data that have been adjusted to standard conditions through use of a computer program, very little difference from the results shown in Table 6-I is observed. The adjusted data show all engine thrust levels to be within 0.40 percent of those achieved during vehicle acceptance test.

Three minor engine performance shifts were observed during S-II burn. Engine No. 1 experienced two performance increases, each approximately 6672 Newtons (1500 Ibf), during the first 35 seconds of mainstage operation.

Table 6-I. S-II Engine Performance Deviations (ESC +61 Seconds)

PERCENT PERCENT INDIVIDUAL AVERAGE PARAMETER ENGINE PREDICTED RECONSTRUCTED DEVIATION DEVIATION

Thrust, 1,034,683 (232,606) 1,035,083 (232,696) 0.04 1,016,663 (228,555) 1,017,517 (228,747) 0.08 Newtons 1,023,514 (230_095) 1,017,228 (228,682) -0.61 -0.20 (Ibf) 1,042,085 (234,270) 1,039,576 (233,706) -0.24 1,034,665 (232,602) 1,032,112 (232,028) -0.25

Specific 4173.7 (425.6) 4169.8 (425.2) -0.09 Impulse 4159.0 (424.1) 4170.8 (425.3) 0.28 N-s/kg 4175.7 (425.8) 4165.9 (424.8) -O.23 -0.06 (Ibf-s/Ibm) 4155.1 (423.7) 4157.0 (423.9) 0.05 4175.7 (425.8) 4162.9 (424.5) -0.30

Engine Flowrate 247.9 (546.6) 248.2 (547.2) 0.II 244.5 (539.0) 244.0 (537.9) -0.20 -0.39 -0.14 kg/s 245.1 (540.4) 244.2 (538.3) (Ibm/s) 250.8 (552.9) 250.I (551.3) -0.29 247.8 (546.3) 240.0 (546.7) 0.07

Engine Mixture 1 5.57 5.57 0 Ratio 2 5.56 5.55 -0.18 LOX/Fuel 3 5.59 5.57 -0.36 0.29 4 5.53 5.54 0.18 5 5.49 5.59 1.82

NOTE: Values exclude pressurization flow.

6-7 A thrust decrease of about the same magnitude occurred in engine No. 2 after 64 seconds of mainstage operation. These shifts are indicative of changes in the Gas Generator (GG) oxidizer system flow resistance and are not considered detrimental to engine operation.

Amplified main chamber pressure processed with a 25 hertz low pass filter revealed no high amplitude, low frequency oscillations as experienced on AS-503 and AS-504. As in the flight of AS-505, CECOprecluded any oscillation buildup. 6.4 S-II SHUTDOWNTRANSIENTPERFORMANCE

Engine shutdown sequence was initiated by the stage LOX low level sensors. The LOX depletion cutoff system again included a 1.5-second delay timer. As in the AS-504 and AS-505 flights, this resulted in engine performance decay prior to receipt of the cutoff signal. Due to early CECOhowever, the precutoff decay was greatly reduced compared to AS-504 without CECO. Only engine No. 1 exhibited a significant thrust chamber pressure decay, decreasing 77.9 N/cm2 (113 psi) in the final 0.25 second before cutoff. All other outboard engines thrust chamber pressure decays were of the order of 20.7 N/cm2 (30 psi).

At OECOsignal (548.22 seconds), total vehicle thrust was down to 2,783,479 Newtons (625,751 Ibf). _ Vehicle thrust dropped to 5 percent of this level within 0.75 second. The stage cutoff impulse through the 5 percent thrust level was estimated to be 581,916 N-s (130,820 Ibf-s). No unusual features were apparent in the center engine thrust decay data following CECO,with the decay to 5 percent thrust occurring in approximately 0.3 second. 6.5 S-II STAGEPROPELLANTMANAGEMENT

The propellant management system performed satisfactorily during the propellant loading operation and during flight. The S-II stage employed an open-loop system utilizing fixed, open-loop commandsfrom the IU rather than feedback signals from the tank mass sensing probes. (Open-loop operation was also used on AS-503 and AS-505. It is also planned for use on all subsequent vehicles.)

The facility Propellant Tanking Control System (PTCS) and the propellant management system successfully accomplished S-II loading and replenishment. During the prelaunch countdown, all propellant management subsystems operated properly with no problems noted.

Open-loop PU system operation commencedwhen "High EMRselect" was commanded at ESC+5.5 seconds, as planned. The PU valves then moved to the high EMR position, providing a nominal high EMRof 5.50 for the first phase of Programed Mixture Ratio (PMR). The IU commandto shift EMRfrom high to low was initiated at ESC+331.8 seconds (6 seconds later than predicted) upon attainment of a preprogramed characteristic velocity as sensed by the LVDC. Approximately 5.5 seconds of this deviation is attributed to improper scaling in the inflight calculations of velocity within the LVDC(refer to

6-8 paragraph 10.2.1), and the remainder is due to variations between the actual and predicted flight performance. The IU commandcaused the PU valves to be driven to the low EMRposition, providing an average EMRof 4.34 (versus a predicted average EMRof 4.33) for the low mixture ratio portion of the flight. The actual shift from high to low EMRoccurred 9.5 seconds late when com- ,pared with the final propulsion prediction. The additional 3.5 seconds result! from a propulsion and characteristic velocity presetting mismatch that was known prior to flight.

Engine No. 3 PU valve position monitor exhibited erratic characteristics during the S-IC and S-II boost operational periods. Analysis of the limited measurements available did not reveal any PU computer, telemetry or engine malfunction. The PU valve telemetry potentiometer is the most likely cause of this problem.

The open-loop PU control system responded as expected during flight and no instabilities were noted. The open-loop PU error at OECOwas approximately +567 kilograms (+1250 Ibm) LH2 versus a 3-sigma tolerance of +_1134 kilograms (+_2500 Ibm).

Based on PU system data, propellant res%duals (mass in tanks and sumps) at OECOwere 816 kilograms (1800 Ibm_LOX, and 2572 kilograms (5671 Ibm) LH2, versus the predicted 657 kilograms (1448 Ibm) LOX and 1966 kilograms (4335 Ibm) LH2. An updated analysis using AS-505 LOX depletion data indicated a higher than predicted LOX residual would occur on AS-506. S-II burn time was reduced approximately 4 seconds and the LH2 residual at OECO was increased 432 kilograms (952 Ibm) due to the late PU valve step time.

Table 6-2 presents a comparison of propellant masses as measured by the PU probes and engine flowmeters. The best estimate propellant mass is based on integration of flowmeter data utilizing the propellant residuals determined from PU system data corrected for nominal tank mismatch at OECO. Best estimates of propellant mass loaded are 370,778 kilograms (817,425 Ibm) LOX, and 71,615 kilograms (157,885 Ibm) LH2 which correlates closely with the postlaunch trajectory simulation. These mass values were 0.24 percent less than predicted for LOX and 0.07 percent less than predicted for LH2. 6.6 S-II PRESSURIZATIONSYSTEMS

6.6.1 S-II Fuel Pressurization System

LH2 tank ullage pressure, actual and predicted, is presented in Figure 6-4 for autosequence, S-IC boost and S-II boost. The LH2 tank vent valves were closed at -96 seconds and the ullage was pressurized to 24.8 N/cm2 (36 psia) in approximately 27 seconds. One makeup cycle was required at -40 seconds as a result of thermal pressure decay. Venting occurred during S-IC boost as anticipated. One venting cycle'was indicated on vent valve No. 1 between 93 and I00 seconds. There was no indication that vent valve No. 2 opened.

6-9 Table 6-2. S-II Propellant Mass History

ENGINE FLOWMETER EVENT INTEGRATION RANGE TIME UIIITS PREDICTED PU SYSTEM ANALYSIS (BEST ESTIMATE) LOX LH2 LOX LH2 LOX LH2 Ground kg 371,672 71,668 371,899 71,718 370,778 71,615 Ignition (I bin) (819,397) (158,000) (819_896) (t58,111) (817,425) (157,885)

S-II ESC kg 371,672 71,668 371,697 71,627 370,778 71,615 (Ibm) (819,397) (158,000) (819,452) _(157,910) (817,425) (157,885)

S-II PU Valve Step kg 38,217 10,751 53,432 13,503 35,884 10,469 (497.60 sec) (Ibm) (84,254) (23,703) (117,797) (29_768) (79,111) (23,080)

S-II OECO kg 657 1966 816 2572 816 2572 (Ibm) (1448) (4335) (1800) (5671) (1800) (5671) S-II Residual At k9 544 1916 730 2531 730 2531 (Ibm) Stage Separation (1199) (4224) (1609) (5579) (1609) (5579)

NOTE: Table is based on mass in tanks and sump only. Propellant trapped external to tanks and LOX sumD is not included.

I I I I I I _7 VENT VALVE NO. 1 OPEN _7 VENT VALVES OPEN ACTUAL PREDICTED _7 VENT VALVE NO. 1 CLOSE _7 VENT VALVE NO. 1 CLOSE (NOTE : PREDICTION COINCIDES _7 S-II ESC _7 S-II OECO, 548.22 WITH DATA) --45 3O I _7 S-II CECO, 460.6 _7 VENT VALVE NO. 2 CLOSE,. _7 LH2 STEP PRESS, 461.6 (POST OECO) I 4O

25 J _35

\M _z9 30 5 2O REDLINE

<, _25 .J \ _ MINIMUM START REQUIREMENT 15

_20

I0 --15 -200 -I00 0 I00 200 300 400 500 600

RANGE TIME, SECONDS

Figure 6-4, S-II Fuel Tank Ullage Pressure

6-10 Differential pressure across the vent valve was kept below the low-mode upper limit of 20.3 N/cm 2 (29.5 psid). Ullage pressure at S-II engine start was 19.6 N/_m2 (28.5 ,psia) meeting the minimum engine start requirement of 18.6 N/cm z (27 psia). The LH 2 tank valves were switched to the high vent mode immediately prior to S-II engine start.

LH2 tank ullage pressure was maintained within the regulator range of 19.7 to 20.7 N/cm2 (28.5 to 30 psia) during burn until the LH 2 tank pressure regulator was stepped open at 461.6 seconds. Ullage pressure increased to 22.1 N/cm2 (32 psia). The LH2 vent valves started venting at 477 seconds and continued venting throughout the remainder of the S-II flight. Ullage pressure remained within the high-mode vent range of 21 to 22.7 N/cm2 (30.5 to 33 psia).

Figure 6-5 shows LH 2 total inlet pressure, temperature and NPSP. The

• z_ parameters were close to predicted values. The NPSP supplied exceeded that required throughout the S-II burn phase of the flight.

6.6.2 S-II LOX Pressurization System

LOX tank ullage pressure, actual and predicted, is presented in Figure 6-6 for autosequence, S-IC boost and S-II burn. After a two-minute cold helium chilldown flow through the LOX tank, the vent valves were closed at -185.3 seconds and the LOX tank was prepressurized to the pressure switch setting of 27.1 N/cm 2 (39.3 psia) in approximately 42 seconds. One pressure makeup cycle was required at -125 seconds as a result of pressure decay, which was followed by the slight pressure increase caused by LH2 tank prepressurization. Ullage pressure was 26.9 N/cm2 (39 psia) at engine start.

The LOX regulator remained at its minimum position until 240 seconds because the ullage pressure was above the regulator range of 24.8 to 26.5 N/cm 2 (36 to 38.5 psia). A slight decrease in ullage pressure prior to LOX regulator step pressurization indicated normal performance of the LOX regulator. LOX step pressurization (261.6 seconds) caused the usual characteristic surge in ullage pressure followed by a slower increase until LOX tank ullage pressure reached a maximum of 28.3 N/cm2 (41 psia) at 383.4 seconds when the No. 1 vent valve cracked. Ullage pressure was 27.9 N/cm 2 (40.5 psia) at CECO. Vent valve No. 1 reseat occurred at 27.9 N/cm2 (40.5 psia) after EMR shift. The LOX tank vent valve No. 2 did not open.

LOX pump total inlet pressure, temperature and NPSP are presented in Figure 6-7. The NPSP supplied exceeded the requirement throughout the S-II boost phase. The total magnitude of LOX liquid stratification was greater than predicted, but was similar to AS-505. The 1.5-second time delay in the LOX low level cutoff circuit makes it very difficult to predict an accurate cutoff temperature.

6-11 S-II ESC LH2 STEP PRESSURIZATION, 461.6 S-ll CECO, 460.6 _7 S-ll OECO

25 I ! ! ACTUAL - 36 ._J E ,r-- PREDICTION - 34 _z 23 F-- F-- LU W WL_ ._jr_ jl.__-- - 32 Z_ if) 21 ...... 30 C_JLU _LO "l-C_ -rr_ __J E__ __Jr_ 19 --28

23 - -419 " o 0

W 22 - -420 r_ _- 421 i,_-i i---_ "'F-- _a< ._J=_ Zr_ Zr'1 - -4_L H ,,, H W 21 r_ r_ -423 m ,,, -r-w _ --J I--- 2O

12 --16

10 -I 4

t'M -12 _ 8 u_ Z ! -I0 ! _ 6 r_ - 8 Z z PSP REQUIREMENT --MINIMUM N - 6

- 4

0 50 I00 150 200 250 300 350 4O0

TIME FROM ESC, SECONDS

i , , , ' 163 263 363 463 563

RANGE TIME, SECONDS

Figure 6-5. S-II Fuel Pump Inlet Conditions

6-12 I I I I _7 S-II ESC _7 S-II CECO ACTUAL _7 LOX STEP PRESSURIZATION S_7 VENT VALVE NO. 1 CLOSE --- - PREDICTION _7 VENT VALVE NO. 1 OPEN S_7 S-II OECO -45 3O I

-40 "_, z J

25; Q -35 c_ L_ c/3 C_ RMINIMUM START REQUIREMENT -30 N J 20 5 _..I

x o -25 o,

--2O

I0 _7 S_7 %_7 -15 -200 -I00 0 I00 200 300 400 5OO 600

RANGE TIME, SECONDS

Figure 6-6. S-II LOX Tank Ullage Pressure

6.7 S-II PNEUMATIC CONTROL PRESSURE SYSTEM

Performance of the stage pneumatic control system was satisfactory. Main receiver pressure and regulator outlet pressure were within predicted limits throughout system operation, Regulator outlet pressure was within the operating band of 476 to 527 N/cm 2 (690 to 765 psia) except during valve actuations which follow S-II ESC, CECO and OECO events. The makeup period for the regulator outlet pressure to return to its operating band after valve closures did not exceed 17 seconds. This is within the normal recovery time.

Pressure decay in the main receiver from facility supply vent at -30 seconds to the initial valve actuation at 168 seconds was negligible. Main receiver pressure was 2086 N/cm 2 (3025 psia) at S-II engine start.

6.8 S-II HELIUM INJECTION SYSTEM

The performance of the helium injection system was satisfactory. Require- ments were met and parameters were in good agreement with predictions. The supply bottle was pressurized to 2137 N/cm2 (3100 psia) prior to liftoff and by ESC was 552 N/cm 2 (800 psia). Helium injection system average total flowrate during supply bottle blowdown (-30 to 163 seconds) was 2.0 SCMM (70.4 SCFM).

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S-IVB PROPULSION

7.1 SUMMARY

The J-2 engine operated satisfactorily throughout the operational phase of first and second burn. Shutdowns for both burns were normal. S-IVB first burn duration was 147.1 seconds which was 3.4 seconds more than predicted. The engine performance during first burn, as determined from standard altitude reconstruction analysis, deviated from the predicted by +0.20 percent for thrust and +0.05 percent for specific impulse. The S-IVB stage first burn Engine Cutoff (ECO) was initiated by the Launch Vehicle Digital Computer (LVDC) at 699.34 seconds.

The Continuous Vent System (CVS) adequately regulated LH 2 tank ullage pressure at 13.4 N/cm2 (19.5 psia) during orbit, and the Oxygen/Hydrogen (02/H2) burner satisfactorily achieved LH2 and LOX tank repressurization for restart.

Engine restart conditions were within specified limits. The restart at full open Propellant Utilization (PU) valve position was successful.

S-IVB second burn duration was 346.9 seconds which was 1.7 seconds less than predicted. The engine performance during second burn, as determined from the standard altitude reconstruction analysis, deviated from the predicted by -0.56 percent for thrust and +0.05 percent for specific impulse. The S-IVB stage second burn ECO was initiated by the LVDC at 10,203.07 seconds.

Subsequent to second burn, the stage propellant tanks were safed satisfactorily, with LOX dump imparting a 17 m/s (55.8 ft/s) velocity change to the stage.

7.2 S-IVB CHILLDOWN AND BUILDUP TRANSIENT PERFORMANCE FOR FIRST BURN

The propellant recirculation system performed satisfactorily, meeting start and run box requirements for fuel and LOX as shown in Figure 7-I.

The thrust chamber temperature at launch was well below the maximum allowable redline limit of 172°K (-150°F). At S-IVB first burn Engine Start Command (ESC), the temperature was 164°K (-164°F), which is within the requirement of 150 ±61.1°K (-189.6 ±IIO°F).

7-I FUEL PUMP INLET PRESSURE, psia 16 20 24 28 32 36 4O 44 25 I -415

TIME FROM FIRST ITEM ESC, SEC 24 1 0 2 20 I -417 3 48 4 ECO o 23

-419_ c_

F- 22 ..... I -42t _ 4

l l (., I 3 (i',',:_2 21 f

EL !. ,, -423 ENGINE START LIMITS I 2O z/ I I,

-425

19 I0 12 14 16 18 20 22 24 26 28 30 32 FUEL PUMP INLET PRESSURE, N/cm 2

LOX PUMP INLET PRESSURE, psia 28 32 36 40 44 48 52 56 I IOO f - 28O TIME FROM FIRST ITEM ESC, SEC

1 0 98 2 4O -284 3 66 4 68 5 ECO I w 96 I -288 cL I <=

k-- 94 I I • / I -292 _ 92 I 251 /, I_ J -296 _ENGINE START LIMITS_ J 9O / l

-300

88 I 18 20 22 24 26 28 30 32 34 36 38 40 LOX PUMP INLET PRESSURE, N/cm 2

Figure 7-I. S-IVB Start Box and Run Requirements- Fi rst Burn

7-2 The chilldown and loading of the engine Gaseous Hydrogen (GH2) start sphere and pneumatic control sphere prior to liftoff were satisfactory. The engine control bottle pressure and temperature at l iftoff were 2124 N/cmL (3080 psia) and i63°K (-i69°F), respectively. At first ESC the start tank conditions were within the required S-IVB region of 896.3 +_68.9 N/cm2 and 133.1 +_44.4°K (1300 +_I00 psia and -220 ±80°F). The discharge was completed and the refill initiated at first burn ESC +3.7 seconds. The refill was satisfactory. The first burn start transient was satisfactory with thrust buildup within the limits set by the engine manufacturer. This buildup was similar to the thrust buildups observed on the AS-501 through AS-505 flights. The PU valve was in proper null position prior to first start. The total impulse from first Start Tank Discharge Valve (STDV) open to STDV+2.5 seconds was 857,243 N-s (192,716 Ibf-s). This was more than the value of 833,615 N-s (187,404 Ibf-s) obtained during the same interval for the acceptance test.

First burn fuel lead generally followed the predicted pattern and resulted in satisfactory conditions as indicated by the thrust chamber and fuel injector temperatures. 7.3 S-IVB MAINSTAGEPERFORMANCEFORFIRST BURN

The propulsion reconstruction analysis showed that the stage performance during mainstage operation was satisfactory. A comparison of predicted and actual performance of thrust, total flowrate, specific impulse, and mixture ratio versus time is shown in Figure 7-2. Table 7-I shows the specific impulse, flowrates and mixture ratio deviations from the predicted at the STDV+137-second time slice when engine performance stabilized. This time slice performance is the standardized altitude performance which is comparable to engine tests. The 137-second time slice performance for first burn thrust was 0.20 percent higher than predicted. Specific impulse performance for first burn was 0.05 percent higher than predicted. S-IVB burn duration was 147.1 seconds which was 3.4 seconds more than predicted.

The helium control system for the J-2 engine performed satisfactorily during first mainstage ooeration. Since the engine bottle was connected with the stage ambient repressurization bottles there was little pressure decay. Approximately 0.19 kilogram (0.42 Ibm) of helium was consumed during first burn.

The PU valve position shifted slightly away from the null position during engine operation. This shift was in the closed (high Engine Mixture Ratio [EMR]) direction and amounted to 0.7 degree during first burn and 0.6 degree during second burn. These shifts are approximately the same as those observed on the AS-505 flight and the S-IVB-508 and S-IVB-509 acceptance tests. Valve position shifts during engine operation have occurred only in engines with PU valves containing rotated baffles. The magnitude

7-3 ENGINE MIXTURE TOTAL FLOWRATE, SPECIFIC IMPULSE, THRUST, 106 N RATIO, LOX/FUEL kg/s N-s/kg

r,o rO --_ PO _ 0 0 0 (.0 C) (D 0 0 o'i o o o o_0 Lo

I I

I I i I I i -0 "0 :I=, I Ill. ' I I mm-_

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..______0 0 0 0 0 0 0 0 0 0 _ 0 0"I _ --' _ 0 _ TOTAL FLOWRATE, SPECIFIC IMPULSE, I bm/s Ibf-s/Ibm THRUST, 106 lbf Table 7-I. S-IVB Steady State Performance - First Burn (STDV +137-Second Time Slice at Standard Altitude Conditions)

FLIGHT PERCENT PARAMETER PREDICTED RECONSTRUCTION DEVIATION DEVIATION FROM PREDICTED

Thrust N 899,399 901,223 1824 (Ibf) (202,193) (202,603) (410) 0.20 Specific Impulse N-s/kg 4202 42O4 2 (Ibf-s/Ibm) (428.5) (428.7) (O.2) 0.05 LOX Flowrate kg/s 177.94 178.24 O. 30 (Ibm/s) (392.30) (392.95) (0.65) 0.17 Fuel Flowrate kg/s 36. O9 36.14 0.05 (Ibm/s) (79.57) (79.67) (0.10) 0.14 Engine Mixture Ratio LOX/Fuel 4.930 4.932 0.002 0.04

of the flow forces for a PU valve with a rotated baffle (determined from recent engine manufacturer testing) combined with the PU electronics gain factor (feedback to control) results in an expected valve displacement of approximately 0.75 degree.

It was concluded that the shift in valve position during the AS-506 flight was due largely to the increased flow forces resulting from the rotated baffle and possibly partly due to an electrical phase change. This observed 0.6 to 0.8 degree shift in valve position during null PU operation is expected to occur on AS-507 and subsequent flights.

7.4 S-IVB SHUTDOWN TRANSIENT PERFORMANCE FOR FIRST BURN

S-IVB ECO was initiated at 699.34 seconds by a guidance velocity cutoff command. The ECO transient was satisfactory and agreed closely with the acceptance test and predictions. The total cutoff impulse to zero percent of rated thrust was 188,302 N-s (42,332 Ibf-s). Cutoff occurred with the PU valve in the nulJ position.

7.5 S-IVB PARKING ORBIT COAST PHASE CONDITIONING

The LH2 CVS performed satisfactorily, maintaining the fuel tank ullage pressure at an average level of 13.4 N/cm2 (19.5 psia).

The continuous vent regulator was activated at 758.5 seconds. Continuous venting was terminated at 9320.4 seconds. The CVS performance is shown in Figure 7-3.

7-5 _FIRST ECO 2O E (.)

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> 0 0 I000 2000 3000 4000 5OOO RANGE TIME, SECONDS ,, V I I I 0:00:00 0:30:00 I:00:00 1:30:00 RANGE TIME, HOURS:MINUTES:SECONDS Figure 7-3. S-IVB CVS Performance - Coast Phase (Sheet 1 of 2)

7-6 S_TSTART OF T6 S_TSECOND ESC 2O

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z 15 20

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0 0 > 5000 6000 7000 8000 9000 I0,000 RANGE TIME, SECONDS

,,, _ . , l , n '_;7 _7 1:30:00 2:00:00 2:30:00 RANGE TIME, HOURS:MINUTES:SECONDS

Figure 7-3. S-IVB CVS Performance - Coast Phase (Sheet 2 of 2)

7-7 Calculations based on estimated temperatures indicate that the mass vented during parking orbit was 966 kilograms (2130 Ibm) and that the boiloff mass was 1081 kilograms (2383 Ibm). 7.6 S-IVB CHILLDOWNANDRESTARTFORSECONDBURN

Repressurization of the LOX and LH2 tanks was satisfactorily accomplished by the 02/H2 burner. Helium heater "ON" commandwas initiated at 9320.2 seconds. The LH2 repressurization control valves were opened at helium heater "ON" +6.1 seconds and the fuel tank was repressurized from 13.4 to 20.8 N/cm2 (19.5 to 30.2 psia) in 193.7 seconds. There were 12.1 kilograms (26.7 Ibm) of cold helium used to repressurize the LH2 tank. The LOX repressurization control valves were opened at helium heater "ON" +6.3 seconds and the LOX tank was pressurized from 25.0 to 27.8 N/cm2 (36.2 to 40.3 psi) in 145.3 seconds. There were 1.95 kilograms (4.3 Ibm) of helium used to repressurize the LOX tank. LH2 and LOX ullage pressures are shown in Figure 7-4. The burner continued to operate for a total of 454.8 seconds providing nominal propellant settling forces. The performance of the AS-506 02/H2 burner was satisfactory as shown in Figure 7-5.

The engine start sphere was recharged properly and maintained sufficient pressure during coast. Between first and second burns, the rate of pressure increase was less than predicted. Also the start bottle relief valve regulated higher than the nominal setting.

The engine control sphere gas usage was as predicted during the first burn; the ambient helium spheres recharged the control sphere to a nominal level adequate for a proper restart.

The S-IVB propellant recirculation systems performed satisfactorily and provided adequate conditioning of propellants to the J-2 engine for the restart as shown in Figure 7-6. Second burn fuel lead resulted in satisfactory conditions as indicated by the thrust chamber and fuel injector temperatures. The start tank performed satisfactorily during the second burn blowdown and recharge sequence.

The second burn start transient was satisfactory with thrust buildup similar to the thrust buildup on flights AS-501 through AS-505. The PU valve was in the proper full open (4.5 EMR)position prior to the second start.

The total impulse from STDVto STDV+2.5 seconds was 794,114 N-s (178,524 Ibf-s). This was less than the value of 833,615 N-s (187,404 Ibf-s) obtained during the same interval for the acceptance test.

The helium control system performed satisfactorily during second burn mainstage. There was little pressure decay during the burn due to the connection to the stage repressurization system. Approximately 0.553 kilogram (1.22 Ibm) of helium was consumed during second burn.

7-8 S_7 HELIUM HEATER ON, 9320.2 LH2 AND LOX CRYOGENIC REPRESS VALVES OPEN, 9326.3 and 9326.5 S_TTERMINATION OF LOX TANK REPRESS _7 TERMINATION OF LH2 TANK REPRESS S_7 HELIUM HEATER OFF 35 5O

N E f_ Z 30

:::D c_ v') L.U or) cz) 4O n,l r_ r_

a,i

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20 -r_ ._./ --r-

16 I0 -I00 0 1O0 200 300 400 5O0 TIME FROM HELIUM HEATER ON, SECONDS

| , V , , ! 02 : 34 : O0 02:36:00 02:38:00 02:40:00 02:42:00 02:44:00 RANGE TIME, FIOURS:MINUTES:SECONDS

Figure 7-4. S-IVB Ullage Conditions During Repressurization Using 02/H 2 Burner

7-9 HELIUMHEATERON, 9320.2 LH2ANDLOXCRYOGENICREPRESSVALVESOPEN,9326.3 AND9326.5 S_TTERMINATIONOFLOXTANKREPRESS S_7TERMINATIONOFLH2TANKREPRESS S_7HELIUMHEATEROFF 0.I0 - 0.20 0 O. 08 0 0.15 ,,, 0.06

0 ..---___.._ 0 Z -0.I0 _ 0.04 Z =E: ,,=:Z =Z) F--

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0 O. 020

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E rY__3 -_r-- 0 -- _ 0.010 0 LL._/ 002. ,,_ :_,g Z Z -- x O. 005 _0 --Ix -r._j _..10 \ -r" -.1 0 200 -40 4- Z ..Q 150 F-- c_ _2 v') r_ -I- F- I00 -20 i r_ I--- i,i z L..U 50 Z tY" r_ L EE) __ -0 -I00 0 I00 200 300 400 50O TIME FROM HELIUM HEATER ON, SECONDS g7 , _7 , _7, _7 , , _ , 02:34:00 02:36:00 02:38:00 02:40:00 02:42:00 02:44:00 RANGE TIME, HOURS:MINUTES:SECONDS

Figure 7-5. S-IVB 02/H 2 Burner Thrust and Pressurant Fl0wrates

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The 172-second time slice performance for second burn thrust was 0.56 percent lower than predicted. Specific impulse performance for second burn was 0.05 percent higher than predicted. A shift in performance at the null PU valve position (-1.5 degrees) occurred during second burn. A shift in the Gas Generator (GG) system resistance is suspected as being the cause of the down shift of 6859 Newtons (1542 Ibf). Also, during second burn several PU valve system resistance shifts are believed to have occurred.

S-IVB second burn duration was 346.9 seconds which was 1.7 seconds less than predicted.

7.8 S-IVB SHUTDOWN TRANSIENT PERFORMANCE FOR SECOND BURN

S-IVB ECO was initiated at 10,203.07 seconds by a guidance velocity cutoff command which resulted in 1.70 seconds shorter than predicted second burn time. The transient was satisfactory and agreed closely with the acceptance test and predictions. The total cutoff impulse to zero percent of rated thrust was 239,061 N-s (53,743 Ibf-s). Cutoff occurred with the PU valve in the null position.

7.9 S-IVB STAGE PROPELLANT MANAGEMENT

The PU system was operated in the open-loop mode. The PU system successfully accomplished the requirements associated with propellant loading.

A comparison of propellant mass values at critical flight events, as determined by various analyses, is presented in Table 7-3. The best estimate full load propellant masses were 0.25 percent greater for LOX and 0.25 percent greater for LH2 than the predicted values. These deviations were well within the required loading accuracies.

Extrapolation of propellant level sensor data to depletion, using propellant flowrates, indicated that a LOX depletion cutoff would have occurred approximately 12.4 seconds after second burn velocity cutoff.

During first burn, the PU valve was positioned at null for start and remained there, as programed, for the duration of the burn. The PU valve was commanded to the 4.5 EMR position 119.9 seconds prior to second burn start command, and remained there for 246.1 seconds. At second ESC +126.2

7-12 ACTUAL %_7 SECONDESC _7 SECOND ESC ACTUAL PREDICTED _7 SECONDECO PREDICTED _7 SECOND ECO PREDICTED BAND PREDICTED BAND 220 i.C 0.22 - 480

- 470

210 O. 9 0.20 - 460 .g i _c I-.- 0.18 _- 0.8 0 190 _-- - 420 o_ -410 _

- 400 0.16 ;.... C.7 180 - 390

43CD 380 170

- 435 5.2

I _ _25D I ...._ .... c_ z c_ 4.9 J 430 uo _200

_ 4.6

I 425

_ w 4150 _o 4.3 _J__

! 420 4.0 _I00 300 350 5O 1O0 150 200 250 300 350" 0 50 1O0 150 200 250 TIME FROM STDV + 2.5 SECONDS TIME FROM STDV +2.5 SECONDS

I I I I I I I I I 2:46:00 2:48:00 2:50:00 2:44:00 2:46:00 2:48:00 2:50:00 2:44:00 RANGETIME, HOURS:MINUTES:SECONDS RANGE TIME, HOURS:MINUTES:SECONDS

Figure 7-7. S-IVB Steady-State Performance - Second Burn Tabl e 7-2. S-IVB Steady State Performance - Second Burn (STDV +172-Second Time Slice at Standard Altitude Conditions)

PERCENT PARAMETER PREDICTED SECOND BURN FLIGHT RECONSTRUCTION DEVIATION DEVIATION i FROM PREDICTED Thrust N 899,399 894,364 -5035 (Ibf) (202,193) (201,061) (-1132) -0.56 Specific Impulse N-s/kg 4202 42O4 2 (I bf-s/Ibm) (428.5) (428.7) (0.2) 0.05 LOX Flowrate kg/s 177.94 176.86 -1.1 (Ibm/s) (392.30) (389.90) ( -2.4) -0.61 Fuel Flowrate kg/s 36.09 35.88 (Ibm/s) (79.57) (79.10) -0.59 Engine Mixture Ratio LOX/Fuel 4.930 4.929 -0.001 -0.02

Table 7-3. S-IVB Stage Propellant Mass History

PU INDICATED EVENT UNITS PREDICTED (CORRECTED) PU VOLUMETRIC FLOW INTEGRAL BEST ESTIMATE

LOX LH 2 LOX LH 2 LOX LH 2 LOX LH 2 LOX LH 2

S-IC Ignition kg 87,100 19,731 87,187 19,761 87,360 19,791 87,119 19,753 87,315 19,780 (Ibm) (192,023) (43,500) (192,215) (43,565) (192,596) (43,631) (192,065) (43,548) (192,497) (43,608)

First S-IVB 87,187 19,756 87,360 19,786 87,119 19,753 87,315 19,757 Ignition (Ib_I (192,023)87,100 (43,500)19,731 (192,215) (43,555) (192,596) (43,621) (192,065) (43,548) (192,497) (43,557) First S-IVB 61,242 14,284 Cutoff (Ib_I 61,539 14,556 61,354 14,380 61,007 14,437 61,300 14,395 (135,670) (32,091) (135,016) (31,491) (135,262) (31,702) (134,497) _31,829) (135,144) (31,736)

Second S-IVB kg 61,406 13,283 61,124 13,207 61,236 13,303 60,884 13,326 61,151 13,301 Ignition (Ibm) (135,377) (29,284) (134,756) (29,116) (135,002) (29,327) (134,227) (29,378) (134,817) (29,324)

Second S-IVB kg 2371 926 2489 974 2484 968 2441 948 2488 970 Cutoff (Ibm) (5228) (2043) (5487) (2147) (5477) (2133) (5381) (2089) (5486) (2139)

seconds the valve was commanded to the null position (approximately 5.0 EMR) and remained there throughout the remainder of the flight.

7.10 S-IVB PRESSURIZATION SYSTEM

7.10.1 S-IVB Fuel Pressurization System

The LH2 pressurization system operationally met all engine performance requirements during prepressurization, boost, first burn, coast phase, and second burn.

7-14 Following the termination of prepressurization, the ullage pressure reached relief conditions, approximately 22.0 N/cm2 (32.0 psia) and remained at that level until just after liftoff as shown in Figure 7-8. A small ullage collapse occurred during the first 5 seconds of boost, and then returned to the relief level at 70 seconds due to self pressurization. All during the burn the ullage pressure was at the relief level, as predicted.

The LH2 ullage pressure was 21.4 N/cm 2 (31.0 psia) at second burn ESC as shown in Figure 7-9. Significant venting during second burn occurred at second ESC +280 seconds when step pressurization was initiated. This behavior was as predicted.

The LH 2 pump inlet Net Positive Suction Pressure (NPSP) was calculated from the pump interface temperature and total pressure. Throughout the burn, the NPSP had satisfactory agreement with the predicted. Figures 7-10 and 7-11 summarize the fuel pump inlet conditions for first and second burns, respectively.

_71NITIATION OF PREPRESSURIZATION S_7FIRST ECO 3O

L) "_ 25 z

30 u_ c_ | ,,, 1,1 _- CTED BAND ,., LJJ C9 c_ | ...._ _j =_15 _ -J _ _ .... _;--2 .... -- _:--__:; =--_I r_--__ 20

i -I 0 1 2 3 4 5 6 7 8 9 RANGE TIME, I000 SECONDS

, , , , I 0:00:00 0:30:00 I:00:00 1:30:00 2:00:00 2:30:00 RANGE TIME, HOURS:MINUTES:SECONDS

Figure 7-8. S-IVB LH 2 Ullage Pressure - First Burn and Parking Orbit

7-15 25 i_7 CRYOGENIC" REPRESSURIZATION

/I. ,. S_7SECOND ECO, 10,203.07 l S_7OPEN CVS, 10,203.8 i .30 CLOSE LATCHING NPV, II,103.1 2O I E it,' , S_7 CLOSE CVS, II,104.1 z I Ii _7 OPEN LATCHING NPV I I I,I _7 CLOSE LATCHING NPV _" 15 I _- OPEN CVS ' _"

(.9 i,i r_ r_ Ill\.| I,i/f ' ,//r- ...... J_'#" .---.j--_m.20 __ i,i I0 .

I @ -- 5

f" / I,-/" X J="_._ _/"_" PREDI I CTED BAND "J

_s/ \I " 0 0 9 II 13 15 17 19

I I _ oooI_ 7 I _ I 2:30:00 3:00:00 3:30:00 4:00:00 4:30;00 5:00:00

2O i I I S_70PEN LATCHING NPV

E I PREDICTED' BAND _15 Z 20 _-

i,I "7 i,i Pc"

L/3 (j') LIJ "' 10 c_

I.a.J i,i (..g lO , _J _J

5 c_ -'r" -r- ...J .._J

0 ...... 19 21 23 25 27 29 RANGE TIME, I000 SECONDS

I I I I I 5:30:00 6:30:00 7:30:00 RANGE TIME, HOURS:MINUTES:SECONDS

Figure 7'9. S-IVB LH2 Ullage Pressure - Second Burn and Translunar Coast

7-16 LL-L

u,_n_1%S,_.L_-I - SUOL%.LpU03 %alUl duJnd Lan_-I £]AI-S "OL-L aan6.L_-I

$0N033S '3Hll 39NVU O7L OOL 089 099 0_9 0_9 009 OBS. 09S I I I I I I l l z_

$0N033S '3S3 IS_I3 NO_3 3NIl 09L O_L O_L OOL OB 09 O_ 08 0 r-- r- 08 "-r- .-i- I"O

--o -o r-" 8_- - t8 "13 i--i z z I-" r- l-rl r-rl -H --I 88 --4 --9 0_tT- - m -o l-rl rrl ;;o 3:: --I _8 ---4 c: c po 8L_- - po rrl m

o o -rl ¸ 178

£L _r _-4F- z-1- _-I _- r--Po r_

¢-- 08- O_ po-o rTq m 8_- 000 oo0 r----4 pO ;_== r-_ Frl r-- 9£:- S_ z (./i 0_- cb I-o 0£

! 0

r I _- I-o Z z -o OL- (_ L -13 -CJ z I,n 0L ,..1° lVnlDv j

0_- £L 033 .LS_]I-Iz_ z'zss 'N3dO AOIS N_Jn_]IS_I4 A '%3 IS_JI3Z_ S_7SECOND ESC, 9848.2 _7 SECONDBURN STDV OPEN, 9856.2 S_7 SECONDECO 15 - 2O

Z 10

r_ ACTUAL r_ -I0 V_ r_ Z Z

j'- (M --r- 2E .d ._J ..... _-_ REQUIRED 0 -0 30 E r_ U - 40 r_ Z

25 - 36 _JLLI ,::_ ,-_ /--ACTUAL I"-- C3 _--_ OOr) 0I.._ - 32 I'--" _"_ I-.- 1._ I,JJ IJ..I r", r'_ 2O :E "" - 28 t.-, i_,,. _-I'-- I..iJ I.LI 0.,,I ,.,....I OJ.--I _NOMINAL PREDICTED - 24 -r- z

15 ,. i J i

v' 24 I..I.. 0 0

LLI F'F ¢",r" "--I - -418 I-- 23 , , . i , , , I--"

rY r,,," LIJ r'.

LLI I--" 22 - -420

t--" F-- LIJ .--I ..J Z Z 21 ' ' - -422

"1- "-m" .--I 20 -J 5O 1O0 150 200 250 300 350 400 TIME FROM SECONDESC, SECONDS I J [ I i , J 2:44:00 2:46:00 2:48:00 2:50:00 RANGETIME, HOURS:MINUTES:SECONDS

Figure 7-11. S-IVB Fuel Pump Inlet Conditions - Second Burn

7-18 7.10.2 S-IVB LOX Pressurization System

LOX tank prepressurization was initiated at -167.5 seconds and increased the LOX tank ullage pressure from ambient to 28.3 N/cm2 (41.1 psia) within 18.5 seconds as shown in Figure 7-12. Three makeup cycles were required to maintain the LOX tank ullage pressure before the ullage temperature stabilized. At -96 seconds the LOX tank ullage pressure increased from 27.4 to 28.5 N/cm2 (39.8 to 41.4 psia) due to fuel tank prepressurization, LOX tank vent purge, and LOX pressure sense line purge. These conditions plus boiloff caused the vent/relief valve to open, holding the pressure at 28.8 N/cm2 (41.8 psia). The pressure remained at this level until liftoff.

During S-IC boost there was a relatively high rate of ullage pressure decay caused by an acceleration effect and subsequent thermal collapse, the decay necessitated one makeup cycle from the cold helium spheres as shown in Figure 7-12.

_7 LOX TANK PREPRESS INITIATION _7 FIRST ESC, 549.2 S_7 FIRST ECO, 699.34

35 5O

n_ E or- u -_ 30 ___MAXlMUM PREDICTED Q. Z - 4O rY 25 -'K _0 _- ACTUAL _---MINIMUM PREDICTED i,I 1,J rY rY - 3O i,J ,., 20

..J

X × 15 - 2O 0 0 =J

I0 0 1 2 3 4 5 6 7 8 RANGE TIME, I000 SECONDS

I , 0:00:00 0:30:00 1:00:00 1:30:00 2:00:00 2:30:00 RANGE TIME, HOURS:MINUTES:SECONDS

Figure 7-12. S-IVB LOX Tank Ullage Pressure - First Burn and Parking Orbit

7-19 One makeup cycle was also required during S-II boost. Although ullage cooling continued during this period, the major cause of the decay again appears to be response to the vehicle acceleration. The LOX tank ullage pressure was 27.7 N/cm2 (40.2 psia) at ESC.

During S-IVB first burn, three over-control cycles were initiated, as predicted. Heat exchanger performance during first burn was satisfactory.

During the coast period between first and second burns the LOX ullage pressure decreased from 29.0 to 25.0 N/cm2 (42.1 to 36.2 psia) which was approximately 5 percent below the predicted minimum. Although this decay was not a problem, it was greater than usual. The ullage pressure decay could have been the result of a combination of factors, including bulkhead heat transfer rate, initial coast ullage temperature, localized boiling rates, and perturbations of the stage. The above possibilities are still under investigation. The decay could also have been the result of leakage through the LOX vent system although a leak of this magnitude could not be detected by stage instrumentation, this possibility cannot be completely eliminated.

Repressurization of the LOX tank prior to second burn was required. The tank ullage pressure was increased from 25.0 to 27.9 N/cm2 (36.2 to 40.4 psia) prior to second ESC. At ESCthe pressure was 27.7 N/cm2 (40.2 psia) satisfying engine start requirements as shown in Figure 7-13.

Pressurization system performance during second burn was satisfactory, having the same characteristics noted during first burn. As predicted, there were no over-control cycles. Heat exchanger performance was satisfactory.

The LOX NPSPcalculated at the interface was 16.81 N/cm2 (24.38 psid) at first burn STDVopen. The minimum NPSPduring burn was 17.1 N/cm2 (24.8 psid) at I00 seconds after ESC. This was 11.4 N/cm2 (16.6 psid) above the required NPSPat that time.

The LOX pump static interface pressure during first burn followed the cyclic trends of the LOX tank ullage pressure. The NPSPcalculated at the engine interface was 16.02 N/cm2 (23.24 psid) at second burn ESC. At all times during second burn, NPSPwas above the required level. Figure 7-14 and 7-15 summarize the LOX pump conditions for the first burn and second burn, respectively.

The cold helium supply was adequate to meet all flight requirements. At first burn ESCthe cold helium spheres contained 171 kilograms (378 Ibm) of helium. At the end of the first burn, the helium mass had decreased to 147 kilograms (325 Ibm). At second burn ESCthe spheres contained 132 kilograms (292 Ibm) of helium. At the end of second burn the helium mass had decreased to 75 kilograms (166 Ibm). Figure 7-16 shows helium supply pressure history.

7-20 SECOND ESC SECOND ECO, 10,203.07 _TLOX TANK NONPROPULSIVE VENT OPENED, 10,204.0 S_TLOX TANK NONPROPULSIVE VENT CLOSED, 10,354.0 _TMANEUVER TO TRANSPOSITION, DOCKING, AND EXTRACTION _/CSM/S-IVB SEPARATION S_TSTART LOX DUMP, 18,187.6 S_'END LOX DUMP, 18,295.8 _TLOX TANK NONPROPULSIVE VENT OPENED, 18,490.8 35 5O

3O 4O

25 m_ ,._ e_ r,.., ACTUAL f- MAXIMUM _REDICTED 30 m 20 u.l -_%.1 _N 15 1If" 2o

× 10 MINIMUM PREDICTED ._.1 lO

0 9 I0 II 12 13 14 15 16 17 !8 19 RANGE TIME, I000 SECONDS

S_7 I ! ! 2:30:00 3:30:00 4:30:00 5:30:00 RANGE TIME, HOURS:MINUTES:SECONDS

_15 20 r_ LU r_ C_ C_ L_

x MAXIMUM PREDICTED. MINIMUM PREDICTED

0 0 19 21 23 25 27 29 RANGE TIME, I000 SECONDS

! ! ! 5:30:00 6:30:00 7:30:00 RANGE TIME, HOURS:MINUTES:SECONDS

Figure 7-13. S-IVB LOX Tank Ullage Pressure - Second Burn and Transl unar Coast

7-21 LOX NPSP, N/cm2 LOX PUMP INLET LOX PUMPTOTAL INLET TEMPERATURE, °K PRESSURE, N/cm2

o DO PO O0 CO 0 0 _ 0 _ 0 0 .--I° I e- ( o I I l .=_ I --I .--I .--I m _0 _-I o I I 0 _ 0 I ---I O0 0"1 O_ m --I .._ C_ 0 0 m r,-.

0 0 Z 1-11 0 X _'_ m m -tl c- 2N o_ o"1 "0 ! Po I k, r_ "s

CD I I i-'rl 0 I 0 z I ,.._° I ...,.1 ° 0 0 I e_ I I I ,,..I ° I tn "M I

c- I I ! I

I I I I I r_O _ _ h_ hO 0 0 0 0

0 LOX PUMP INLET LOX PUMPTOTAL INLET LOX NPSP, psid TEMPERATURE, °F PRESSURE, psia SECOND ESC, 9848.2

SECOND BURN STDV OPEN, 9856.2 SECOND ECO 25

2O -30

E (&

Z 15 20

a_ z 10 Z x REQUIRED X C) 0 .--I ._I \ lO

0 0 35 E I 50 (.J

Z r'-, 3O ,::_ rr' I'-" --'_ I I"-- --'_ OOO 4O O(X'} I--- On -- PREDICTED I--oO I.l.I 25 _-cL.

D-. I-- r_ i-- l.ul Lul X..-I X--.I OZ 20 3O 0 7. 93 --293 I._ 1.--o I---o LU ! _J 92 -294 .--.I z_ ZlJ_l i.-i r_,' I _-._c_."

r_ I..- --295 c_ I--- _',_C

r-, w 91 --296 c_-w r_ r'_ X_'- X_Z OL_ OL_J •-J t--- -.Jl-- --297 90 0 5O 1O0 150 200 250 300 350 4O0 TIME FROM SECOND ESC, SECONDS

i , ,,S 7 2:44:00 2:46:00 2:48:00 2:50:00 RANGE TIME, HOURS:MINUTES:SECONDS

Figure 7-15. S-IVB LOX Pump Inlet Conditions - Second Burn

7-23 _7 FIRST ECO _7 START CRYOGENIC REPRESS

_7 STOP CRYOGENIC REPRESS, 9626.3 X_7 SECOND ESC, 9848.2 _7 SECOND ECO, 10,203.07 _7 START COLD HELIUM DUMP, 10,264 END COLD HELIUM DUMP, 2500

3000 2000 _

1500 n_ E u 2000 C)_ z

W C_

cz) Cz_ W C_ r_ 1000

I000

5OO

0 0 4 8 12 16 20

RANGE TIME, I000 SECONDS

I I 0:00:00 2:00:00 4:00:00 RANGE TIME, HOURS:MINUTES:SECONDS

Figure 7-16. S-IVB Cold Helium Supply History

7-24 7.11 S-IVB PNEUMATICCONTROLSYSTEM

The pneumatic control and purge system performed satisfactorily during all phases of the mission. System performance was nominal during boost and first burn operations. 7.12 S-IVB AUXILIARY PROPULSIONSYSTEM

The Auxiliary Propulsion System (APS) pressurization system demonstrated nominal performance throughout the flight and met control system demands as required.

System pressures and propellant temperatures are presented in Table 7-4, All APS engines performed satisfactorily with chamber pressures ranging from 62 to 69 N/cm2 (90 to I00 psia).

The APS ullage engines were turned on at approximately 700 seconds and 9775 seconds for propellant settling and were turned on a third time at approximately 20,268 seconds to provide additional impulse for the slingshot maneuver.

The propellant consumption curves and predictions are presented in Figure 7-17. Table 7-5 presents the APS oxidizer and fuel consumption at significant events during the flight.

Table 7-4. S-IVB APS Propellant Conditions

PARAMETER MODULE NO. 1 MODULE NO. 2

FUEL OXIDIZER FUEL OXIDIZER

Ullage Pressure N/cm Z 131 to 133 131 to 133 128 to 130 131 to 133 (psia) (190 to 193) (190 to 193) (186 to 188} (190 to 193)

Propellant Manifold Pressure 131 to 132 tl/cm2 133 to 135 133 to 135 130 to 131 (190 to 192) (psia) (193 to 196) (193 to 196) (188 to !go)

Propell ant Temperature (Control Module) 302 to 315 o K 297 to 304 300 to 309 303 to 315 (°F) (75 to 87) (80 to 96) (86 to 107) (84 to 107)

Regulator Outlet Pressure N/cm 2 128.2 to 134.4 128.2 to 134.4 134 to 135 134 to 135 (psia) (186 to 195) (186 to 195) (194 to 196) (194 to 196)

7-25 _7 FIRST ESC, 549.2 1_7 FIRST ECO, 699,34 S_7 SECOND ESC, 9848.2 _7 SECOND ECO, 10,203.07 1_7 LOX DUMP _7 APS ULLAGE BURN

I00 MODULE NO. 1 I00, . MODULE NO. 2 II 90 -200 90 - 200

8o -175 80 - 175

70 CTU 150 70 150 6o _ _ 60 PREDICTED BAND_]'____ 125 _ PREDICTED BAND_ _ I 125 OXIDIZER == 5 --,._ _ <_ =<

I J I00 f_ Oh 40 '_,..__ L ___ ,.-H __ 40

c:c 75 a. _" 75 PREDICTED BAND_

30 PREDICTED BAND--_ /FUEL 30

-5O ---_ _ 50 20 20

I0 -25 I0 25

0 0 0 0 0 8000 16,000 24,000 32,000 0 8000 16,000 24,000 32,000 RANGE TIME, SECONDS RANGE TIME, SECONDS _7 _7 , _z, , ,_,, , ,, , , ,7 O:OO:O0 2:00:00 4:00:00 6:00:00 8:00:00 I0:00:00 O:O0:O0 2:00:00 4:00:00 6:00:00 8:00:00 I0:00:00 RANGE TIME, HOURS:MINUTES:SECONDS RANGE TIME, HOURS:MINUTES:SECONDS

Figure 7-17. S-IVB APS Propellants Remaining Versus Range Time, Module No. 1 and Module No. 2 Table 7-5. S-IVB APS Propellant Consumption

MODULE AT POSITION I MODULE AT POSITION III TIME PERIOD OXIDIZER FUEL OXIDIZER FUEL

KG (LBM) KG (LBr.I) KG (LBM) KG (LBM)

Initial Load 92.16 (203.00) 57.20 (126.00) 92.08 (203.00) 57.20 (126.00)

First Burn 0.43 (0.95) 0.35 (0.76) 0.43 (0.95) 0.33 (0.72) (Roll Control)

ECO to End of 7.30 (16.07) 5.52 (12.18) 7.27 (16.02) 5.53 (12.17) First APS Ullaging

End of First Ullage 4.48 (9.88) 2.77 (6.11) 2.38 (5.24) 1.48 (3.27) Burn to Start of T 6

T 6 to Start of 0.83 (1.83) 0.52 (1.15) 0.12 (0.26) 0.07 (0.16) Second Ullage

Second Ullage Burn 6.12 (13.50) 4.80 (IO.58) 5.57 (12.27) 4.34 (9.57)

Second Burn 0.59 (I.29) 0.38 (0.84) 0.II (o.25) o.07 (0.16) (Roll Control)

ECO to LOX Dump 4.03 (8.88) 2.52 (5.55) 7.55 (16.64) 4.72 (io.41)

LOX Dump 1.27 (2.80) 0.79 (1.74) 3.18 (7.00) 1.98 (4.37)

LOX Dump to Third 1.78 (3.92) I.II (2.45) 1.45 (3.20) 0.91 (2.00) Ullage Burn

Third Ullage Burn 18.06 (39.82) 14.45 (31.85) 19.09 (42.09) 15.09 (33.27)

Third Ullage Burn 4.60 (10.14) 2.97 (6.55) 4.27 (9.42) 2.67 (5.89) to Loss of Data

Total Usage 49.52 (I09.08) 36.16 (79.72) 51.41 (113.34) 37.19 (81.99)

7.13 S-IVB ORBITAL SAFING OPERATIONS

7.13.1 Fuel Tank Safing

The LH2 tank was satisfactorily safed using three programed vent cycles utilizing both the Non Propulsive Vent (NPV) and CVS as indicated in Figure 7-18. The LH2 tank ullage pressure during safing is shown in Figure 7-9. At second ECO, the LH2 tank ullage pressure was 22.4 N/cm2 (32.4 psia) and after three vents had decayed to approximately zero. The mass of GH2 and LH2 vented agrees well with the 1174 kilograms (2589 Ibm) of liquid residual and pressurant in the tank at the end of powered flight.

7.13.2 LOX Tank Dump and Safing

Immediately following second burn cutoff, a programed 150-second vent cycle reduced LOX tank ullage pressure from 27.2 to 13.0 N/cm2 (39.4 to 18.9 psia) as shown in Figure 7-13. Data levels were as expected with 32 kilograms (71 Ibm) of helium and 50 kilograms (II0 Ibm) of GOX being vented overboard. As indicated in Figure 7-13, the ullage pressure then increased due to self-pressurization and sloshing to 16.6 N/cm L (24.1 psia) at initiation of LOX dump.

7-27 LH2 TANK CVS OPEN

LOX TANK NPV VALVE OPEN I

LH2 TANK LATCH NPV OPEN

J-2 ENGINE START SPHEREDUMP J I COLD HELIUM DUMP I I AUX HYDRAULIC PUMPON

STAGE PNEUMATLCCONTROLSPHEREDUMP I -.j APS ULLAGE ENGINES ON I

O0 AMBIENT REPRESSDUMP

J-2 ENGINE LOX DOMEPURGE

LOX DUMP

J-2 ENGINE CONTROLSPHEREDUMP

! | ,A, , , ' ' _' l'z ' '9 ' I0 II 13 I'4 15 1 RANGETIME, I000 SECONDS I '_v ' _, ' ' 2:50:00 3:25:00 3:50:00 4:50:00 5:25:00 RANGETIME, HOURS:MINUTES:SECONDS

Figure 7-18. S-IVB LOX Dump and Orbital Safing Sequence The 108-second LOX tank dump was initiated at 18,187.6 seconds and was satisfactorily accomplished. A steady-state liquid flow of 0.024 m3/s (380.6 gpm) was reached within 50 seconds.

Approximately 79 seconds after dump initiation, the measured LOX flowrate showed a sudden increase indicating that gas ingestion had begun. Shortly thereafter, the LOX ullage pressure began decreasing at a greater rate. Calculations indicate the LOX residual, approximately 921 kilograms (2030 Ibm), was essentially dumped within I00 seconds. Ullage gases continued to be dumped until the programed termination. The tank pressure had decayed to 15.4 N/cm2 (22.4 psig) at this time.

LOX dump ended at 18,295.8 seconds as scheduled by closure of the Main Oxidizer Valve (MOV). A steady-state LOX dump thrust of 4003 Newtons (900 Ibf) was obtained. The total impulse before MOV closure was 318,048 N-s (71,500 Ibf-s), resulting in a calculated velocity increase of 17 m/s (55,8 ft/s). Figure 7-19 shows the LOX flowrate during dump and the mass of liquid and gas in the oxidizer tank. This figure also shows LOX ullage pressure and the LOX dump thrust produced. The predicted curves provided for the LOX flowrate and dump thrust correspond to the quantity of LOX dumped and the actual ullage pressure.

At 195 seconds after the end of LOX dump the LOX NPV valve was opened for the duration of the mission. LOX tank ullage pressure decayed from 15.5 N/cm2 (22.5 psia) at 18,490.8 seconds to zero pressure at approximately 24,000 seconds.

7.13.3 Cold Helium Dump

Cold helium was dumped through the 02/H2 burner LH 2 heating coils and into the LH2 tank, and overboard through the tank vents.

Three separate programed dumps totaling 3537 seconds were made starting at 10,264 seconds, as shown in Figure 7-16. During these periods, the pressure decayed from 365 to 17 N/cm2 (530 to 25 psia). Approximately 73.9 kilograms (163 Ibm) of helium were dumped overboard.

7.13.4 Ambient Helium Dump

The ambient helium in the LOX and LH2 repress spheres was dumped through the fuel tank. The 60-second dump was commanded on at 10,204.8 seconds and started at 10,221.6 seconds when the LH2 tank pressure switch dropped out and allowed the repress valve to open, The pressure in the fuel repress spheres decayed from 2124 to 579 N/cm2 (3080 to 840 psia) and 15.6 kilograms (34.4 Ibm) of helium were dumped.

7.13.5 Stage Pneumatic Control Sphere Safing

The stage pneumatic control sphere was safed by initiating the J-2 engine pump purge and flowing helium overboard through the pump seal cavities.

7-29 sqL '±Sn_Hl s/mq[ wq[ _SS'_4 X04

'31V_M073 alnbl7 0 °0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0, 0 0 ,-- CO '4_ _ c_J _ C'q oJ I",,, 0 o") 0 0 co ,0

0 0,

p.,,, i ) oO 0 E O0 :/ ,.

Z Z .0 cO Z X r',, Z I-- 0 j_ // 0 0 _ '--_ /I _1 I c,J I c,O Z ,--" Z °, 0 _o- ,, _-_ Z _-i /_/ 0 i h I ] I °° 0 c,_ I I '" I '" 0 I--- I--. g_ 0 _ 0 "k_.l I I '" I '" d-; . D_ / I °-o_

C4N °_,- t_ 0 t . (M 0

°, m "/f' i P_ D' o 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 _0 I_ C_ 0 0 0 0 0 fO 0J r- 0 0 0 0 0 .0 0 0 0 0 0 LO 0 0 o_ N 00

s16_ 8WO/N '3_nss3_d N 'ISN_HI '3±V'dMO'l_oInbI4 6_ 'SSVW. XO4 39V77N XO] The safing period of 3600 seconds satisfactorily reduced the potential energy in the spheres.

7.13.6 Engine Start Sphere Safing

The engine start sphere was safed during an approximately 148-second period starting at 10,206.3 seconds. Safing was accomplished by open- ing the sphere vent valve. Pressure was decreased from 776 N/cm2 (1125 psia) to zero with 1.78 kilograms (3.93 Ibm) of hydrogen being vented.

7.13.7 Engine Control Sphere Safing

The engine control sphere safing began at 18,505 seconds. The helium control solenoid was energized to flow helium overboard through the engine purge system. The pressure decayed from 1379 to 103.4 N/cm 2 (2000 to 150 psia) and 0.680 kilogram (1.50 Ibm) of helium was vented during the 1300-second safing period.

7-31/7-32

SECTION8

HYDRAULICSYSTEMS

8.1 SUMMARY

The stage hydraulic systems performed satisfactorily on the S-IC, S-II, and first burn and coast phase of the S-IVB stage. During this period all parameters were within specification limits.

The S-IVB hydraulic system pressure exceeded the upper limit by 0.6 percent just after second burn ignition and remained at this level until 202 seconds into the burn. At this time a step decrease in system pressure to a normal operating level occurred. The pressure remained at this level for the remainder of the burn. Other than this minor deviation system performance was nominal and no other problems were noted.

The manufacturer of the S-IVB engine driven hydraulic pump states that the pump has an output pressure "drift-up" characteristic that could account for this excess pressure. The abrupt pressure changes noted during the burn are probably due to frictional hysteresis within the engine driven pump pressure/flow-regulating mechanism. The pump manufacturer does not consider this condition to indicate impending malfunction of the engine driven pump. 8.2 S-IC HYDRAULICSYSTEM

The performance of the S-IC hydraulic system was satisfactory. All servo- actuator supply pressures, and return pressures and temperatures were within required limits. 8.3 S-II HYDRAULICSYSTEM

System steady-state supply pressures during flight ranged from 2400 to 2468 N/cm 2 (3480 to 3580 psia) with steady-state reservoir pressures ranging from 63 to 70 N/cmL (92 to I01 psia). These pressures were well within the predicted ranges of 2275 to 2620 N/cm 2 (3300 to 3800 psia) and 54 to 72 N/cm 2 (78 to 105 psia), respectively. Reservoir volumes at Engine Cutoff (ECO) ranged from 16 to 21 percent, well within the predicted range of 12 to 34 percent. Reservoir fluid temperatures at ECO

8-I ranged from 304 to 319°K (88 to II5°F) compared to a predicted 300 to 328°K (80 to 130°F). The reservoir fluid temperatures and rate of increase of these temperatures compared well with predicted values.

Throughout the flight, all servoactuators responded to commandswith good precision. The maximumdifference between actuator commandand position was 0.2 degree. Forces acting on the actuators were well below a predicted maximumof 84,516 Newtons (19,000 Ibf). The maximumforce in tension was 32,027 Newtons (7200 Ibf) acting on the pitch actuator of engine No. I. The maximum force in compression was 35,586 Newtons (8000 Ibf) acting on the pitch actuator of engine No. I. 8.4 S-IVB HYDRAULICSYSTEM

The S-IVB hydraulic system performance was nominal during S-IC/S-II boost and S-IVB first burn. The supply pressure was nearly constant at 2503 N/cm2 (3630 psia) as compared to an allowable of 2425 to 2527 N/cm2 (3515 to 3665 psia). System flow requirement was provided by the engine driven hydraulic pump during first burn as indicated by a rise in system pressure after ignition and an auxiliary pump motor current draw of 19.5 amperes. Power extraction by the engine driven pump during burn was 3.64 kw (4.88 horsepower).

Engine deflections were nominal during first burn. During orbital coast, two hydraulic system thermal cycles of 48 seconds duration were programed to start at 3300 and 6100 seconds.

The auxiliary hydraulic pump was turned on at 9497.2 seconds during second burn prestart preparations. System operation was normal with output pressure at 2487 N/cmZ (3610 psia) as shown in Figure 8-I. After second ESC at 9848.2 seconds, as the engine driven pump commencedoperation, the system pressure increased to 2542 N/cm2 (3688 psia) which exceeded the upper limit of 2526 N/cm2 (3665 psia) by 0.6 percent. At 10,050 seconds system pressure dropped below the upper limit to 2505 N/cm2 (3632 psia) and remained steady until 10,233.1 seconds when the auxiliary pumpwas turned off. At 10,050 seconds, as the system pressure dropped, the auxiliary pump motor current increased from 20 to 30 amperes indicating that the auxiliary pump assumed an increased share of the hydraulic load. System temperatures, actuator positions and auxiliary pump current loads were normal during the burn and therefore this slight excess in system pressure did not appear to cause any problems.

The pump manufacturer states that the engine driven hydraulic pump has a "drift-up" characteristic which, when combined with uncompensated thermal: expansion in the pump compensator mechanism, makes a rise in output pressure during second burn highly likely. It should be noted that the predicted upper limit of output pressure does not make allowance for this pressure increase. The excessive system pressure after S-IVB second

8-2 150

_AUXILIARY HYDRAULIC PUMP ON •2OO S_TsEcOND ESC _TAUXILIARY HYDRAULIC PUMP OFF 125 SECOND ECO ..... PREDICTED LIMITS _ACTUAL DATA -160 3OOO I I I00 c_ HYDRAULIC SYSTEM - 4000 PRESSURE I •120 _- .-.. _L._ .-_4..--._-, ..... 75 25OO #-_ -4 £.= .._ _ .... {z3 ACCUMULATORGN2 or} 8o_ PRESSURE I _- 50 c_ ' 3OOO 2000 cz_ 4O 25

1500 q .2000 0 1 340 I I fl I i II _HYDRAULIC PUMP INLET_ 1000 i ! !I OIL TEMP. 100 120 Oo 320 I co 8O AFT BATTERY NO. 2 _'CURRENT LOAD .300 80 _ L...... ,I_._._.F-_.._l_--___ RESERVOIR OIL N = 60 m P__ ....r TEMPERATURE ,=,28o '40 w 4O I--

260 2O .... I

240 9250 9450 9650 9850 10,050 10,250 10,450 9250 9450 9650 9850 10,050 10,250 10,450 RANGE TIME, SECONDS RANGE TIME, SECONDS

I ! 2:35:00 2:40:00 2:45:00 2:50:00 2:35:00 2:40:00 2:45:00 2:50:00 RANGE TIME, HOURS:MINUTES:SECONDS RANGETIME, HOURS:MINUTES:SECONDS

Figure 8-I. S-IVB Hydraulic System - Second Burn start is probably due to this effect. The abrupt changes noted during the burn could be due to frictional hysteresis in the pressure/flow- regulating mechanism shown schematically in Figure 8-2. The pump response to a slowly changing demand would then consist of small step changes in output pressure.

The pump manufacturer indicates that the frictional hysteresis may be due to "silting" or entrapment of particulate matter in the small clearance between the compensator spool and sleeve. These components have a lap fit with a clearance of 0.00076 to 0.0010 centimeter (0.0003 to 0.0004 in.) whereas the circulating fluid has nominal 15 micron (0.00059 in.) filtration. This added friction would increase the required force on the spool before a response could occur. It should be noted however, that the existence of silting is not necessary to explain this high friction. The extremely small clearance of the lapped fit between the spool and sleeve, possibly modified by normal wear in proportion to pump life, could be a sufficient explanation. In any case, the pump manufacturer considers that this condition does not indicate impending malfunction of the engine driven pump.

IKE SPRING

YOKEACTUATING PUMPOUTLET CYLINDER

__OUTPUT PRESSURE

--COMPENSATOR VALVE SPOOL

PUMP INPUT SHAFT --COMPENSATOR VALVE SPRING

----PRESSURE ADJUSTMENT

PUMP YOKE _PUMP INLET

Figure 8-2. S-IVB Engine Driven Hydraulic Pump Schematic

8-4 SECTION 9

STRUCTURES

9.1 SUMMARY

The structural loads and dynamic environments experienced by the AS-506 launch vehicle were well within the vehicle structural capability. The longitudinal loads experienced during flight were nominal. The maximum bending moment condition, 3.75 x 106 N-m (33.2 x 106 Ibf-in.), was experienced at 91.5 seconds and was lower than that experienced on any previous flights. The maximum longitudinal loads on the S-IC thrust structure, fuel tank, and intertank were experienced at 135.2 seconds, Center Engine Cutoff (CECO). On all the vehicle structure above the intertank, the maximum longitudinal loads were experienced at 161.6 seconds, Outboard Engine Cutoff (OECO), at the maximum longitudinal acceleration of 3.94 g.

During the S-IC boost phase, low-level (±0.07 g) 4.8-hertz longitudinal oscillations were detected in the Instrument Unit (IU) and peaked at approximately 107 seconds. These oscillations occurred in the first vehicle mode. Except for AS-502, the amplitudes of the oscillations were slightly higher than those observed on other previous flights.

The S-IVB gimbal block longitudinal measurement recorded a small (±0.037 g at 15.5 hertz) oscillation buildup at 252 seconds (S-II boost phase). Similar oscillations were experienced on the AS-505 flight.

Low-frequency longitudinal oscillations similar to those experienced on the AS-505 flight occurred during AS-506 S-IVB first and second burns. The AS-506 first burn peak amplitude I+0.07 g at 19 hertz) was about 20 percent of the AS-505 peak amplitude _0.3 g at 19 hertz). The second burn oscillations peaked at approximately ±0.12 g (13 hertz) at 10,172 seconds.

9.2 TOTAL VEHICLE STRUCTURES EVALUATION

9.2.1 Longitudinal Loads

The AS-506 vehicle liftoff occurred nominally at a steady-state acceleration of approximately 1.2 g. Transients due to thrust buildup and release resulted in a ±0.13 g maximum longitudinal dynamic acceleration measured

9-I at the IU. The slow-release rod forces measured during liftoff are presented in Figure 9-I.

The longitudinal loads that existed at the time of maximumaerodynamic loading (91.5 seconds) are shown in Figure 9-2. The steady-state longitudinal acceleration was 2.34 g, and the corresponding axial loads experienced were nominal.

As shown in Figure 9-2, the maximumlongitudinal loads imposed on the S-IC thrust structure, fuel tank, and intertank occurred at 135.2 seconds (CECO) at a longitudinal acceleration of 3.71 g. The maximumlongitudinal loads imposed on all vehicle structure above the S-IC intertank occurred at 161.6 seconds (OECO)at an acceleration of 3.94 g.

9.2.2 Bending Moments The lateral loads experienced during thrust buildup and release were much lower than design because of the low-level winds experienced during launch. The wind speed at launch was 3.3 m/s (6.4 knots) at the 18.3-meter (60-ft) level. The comparable launch vehicle and spacecraft redline winds were 18.9 m/s (36.8 knots) and 15.4 m/s (30 knots), respectively. mLmL__ AS-502 AS-503 500 INVALID DATA _I -- AS-504 --._ AS-505 AS-506

400

% z -\. \ . _o 3oo -60 o r_ o 2OO - .\\!!, ',,

1 O0 \ ' "\\

0 '\ \\', ,0 0 0.I 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 TIME FROM RELEASE, SECONDS

I I I I I I I I I I 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 o0 1 .I 1.2 RANGE TIME, SECONDS Figure 9-I. Release Rod Force Time History Comparison

9-2 VEHICLE STATION, in.

! ! 40'00 3000 2000 1000

VEHICLE STATION, m 90 80 70 60 50 40 30 20 I0 0 5O t = 91'.5 SEC ACCELERATION = 2.34 g -I0 ...... t = 135.2 SEC ACCELERATION = 3.71 g 4O .... t = 161.6 SEC ACCELERATION = 3.94 g

z 30 q,_

'6,_o o

0 ..,,J ...... 7., r--, --, 20 G 4 x

),,--i x

Figure 9-2. Longitudinal Load at Maximum Bending Moment, CECO and OECO

The inflight winds that existed during the maximum dynamic pressure phase of the flight were low, 9.6 m/s (18.7 knots), at the ll.4-kilometer (37,400 ft) level and were increasing steadily at higher altitudes, as shown in Figure A-I. The maximum bending moments on AS-506 were less than the bending moments experienced on any previous Saturn V vehicle, less than 15 percent of design criteria. As shown in Figure 9-3, the maximum bending moment of 3.75 x 106 N-m (33.2 x 106 Ibf-in.) was imposed on the S-IC LOX tank at 91.5 seconds. Bending moment computations are based upon measured inflight parameter_ such as thrust, gimbal angle, angle-of-attack, dynamic pressure, and accelerations.

9.2.3 Vehicle Dynamic Characteristics

9.2.3.1 Longitudinal Dynamic Characteristics. The predicted first longitudinal mode frequencies were present throughout the AS-506 S-IC boost phase. As shown in Figure 9-4, the measured frequencies agree well with the analytical predictions. The frequencies are determined by spectral analysis using 5-second time slices.

9-3 VEHICLE STATION, in.

3000 2000. I000 VEHICLE STATION, m 90 80 70 60 50 40 30 20 I0 0 i i t = 91.5 SEC 4O = 1.7 DEG O. 08 B AVG= 0.27 DEG

_30 z--- 0.06

3 20 TOTALNORMAL OAO/ 0.04 _j

_I0 J 0.02

Figure 9-3. Maximum Bending Moment Near Max Q

LEGEND

IO ffI_MLT313 rlK_l _IUU_ CO_IMAND"MODU'LE MEASUREDFIRST MODE 0

N -r- 6

(J J LaJ

C)" 0 i,i 4( I,

2O 4O 60 80 I00 120 140 160 RANGETIME, SECONDS

Figure 9-4. First Longitudinal Modal Frequencies During S-IC Powered Flight

9-4 The S-IC CECO and OECO transient responses measured at the Command Module (CM) and the IU are shown in Figure 9-5. The decay of the CECO amplitudes is comparable to previous flights, indicating that the vehicle damping in the first mode is similar. Peak amplitudes of first mode oscillations versus body station for the 135- to 138-second time slice are shown in Figure 9-6. The amplitudes of several measurements on AS-504, AS-505, and AS-506 are shown in this figure as well as a fit of the predicted first vehicle longitudinal mode through the data points.

The most significant vehicle responses during the S-IC stage boost phase were detected by the IU longitudinal (A2-603) measurements and the S-lC intertank lon_itudinal (AI-II8) measurements. As shown in Figure 9-7, oscillations _4.7 to 5.2 hertz) began at approximately 102 seconds, peaked at 107 seconds, and damped by 125 seconds. The peak amplitude measured at the IU was ±0.07 g at 4.8 hertz. Except for AS-502, oscillations in the same frequency band, but at lower amplitudes, have been observed on other previous flights with an amplitude of ±0.05 g measured on AS-505 at 115 seconds. F-I engine chamber pressures in the 4- to 5-hertz region were below the 0.4 N/cm 2 (0.5 psi) noise floor. The observed oscillations were a response of the first longitudinal mode to flight environmental excitations. POGO did not occur during S-IC boost.

During the S-II stage boost phase, a small response buildup was observed by the S-IVB stage gimbal block longitudinal accelerometer (A12-403) at 252 seconds. The AS-506 amplitude peaked at _0.037 g at 15.5 hertz, as shown in Figure 9-8. A similar response was observed during the AS-505 flight, ±0.035 g (15 to 16 hertz), at 293 seconds. The S-II crossbeam amplitude on AS-505 was ±I.0 g. The crossbeam on AS-506 was not instru- mented; however, since the AS-506 oscillations also occur in the high- gain S-II crossbeam mode (15 to 16 hertz), the AS-506 crossbeam response is estimated to have been ±I g (15.5 hertz), which is well below the design limits.

Low-frequency longitudinal oscillations similar to those experienced on AS-505 occurred during AS-506 S-IVB first and second burns. As shown in Figure 9-9, the AS-506 first burn oscillation frequency ranges were identi- cal (17 to 20 hertz); however, the AS-506 peak amplitude (±0.07 g) was about 20 percent of the AS-505 peak amplitude (±0.3 g) at 19 hertz. The AS-506 oscillations were intermittent, recurring at lower amplitudes throughout the remainder of first burn. The LOX suction line inlet measurement reached a maximum of ±0.12 g and showed similar intermittent responses throughout first burn, as shown in Figure 9-10. The data of Figure 9-11 show that the AS-505 and AS-506 peak amplitudes, determined by spectral analysis using 2-second time slices, occurred at the same frequency and near the same time during flight.

9-5 _7 S-IC CECO _7 S-IC OECO COMMANDMODULE COMMANDMODULE 5 5 4 "__v'_A^ A A ^ ^I_AA^^^^ AAAAAI 3 "v VVVV_VVVV_ VVVVV z 2 0 0 DATA DROPOUT F-- ; :C < 1 mY m_ L_J .J 0 l,t

,=C -2 -2

-3 -3

INSTRUMENT UNIT INSTRUMENT rNIT 6 LO I O_

5 i ,

4 i z DATA I o 0 V- B -DROP( pUT- :C AAAAAAAAAA AAAAA VVV¥1 rv v m_ _UVVV Ivvvv_ i,i _J L_ L_

:^ _ A A A i, A A_Am, V VVVV VyVvv _'

_7 -I _7 134 135 136 137 138 139 161 162 163 164 165 RANGE TIME, SECONDS RANGE TIME, SECONDS

Figure 9-5. Longitudinal Acceleration at CM and IU 4000 100 FIRST LONGITUDINAL MODE .[, 5.2 HERTZ (135 TO 138 SEC) C 360C 9O I /

3200 Oral S 8O A0002-603 _

PREDICTED LM I 2800 7O AS-504 / AS-505 A ! c= E AS-506 ii .r- ! =- z I 2400 O C) _- 60 I

I-- I f-- tz_ I I.J_l W ..,..I .=4 (._) PREDICTEDMODE SHAPE ---_I 2000 := 50

LAJ I I I 1600 4O I I I I 1200 I 3O I . I AOOl -I 184 8OO 20 _ E093-II 19 /_' E058-I 18 400 10 1 LE092-117 E057-I 15 _k I dD E083-115 0 , , , I _ J -0.5 O 0.5 ACCELERATION, Gpeak

Figure 9-6. Peak Amplitudes 'of Vehicle First Longitudinal Mode for AS-504, AS-505, and AS-506

9-7 0.08 '1 I I I I I I I 8 LONGITUDINAL ACCELERATION _ INSTRUMENT UNIT (A0002-603) I l I

0.0'

0.0, 6 uJ '/A u- X/ I

P AMPLITUDE i_ 0.02 r,d, "_

UENCY - AJ \i

0.00

0.04

0.03

o.o2

0.01

0.0 4 80 I O0 120 140 RANGE TIME, SECONDS

Figure 9-7. Frequency and Amplitude of Longitudinal Oscillations During S-IC Boost

9-8 • 20 0.04 LONGITUDINAL ACCELERATION - S-IVB GIMBAL BLOCK (A0012-403)

18 /

17-

i-.- o.02

i 16-

FRE[ UENCY • -_ 15-

14"

13.

12 0.0 200.0 220.0 240.0 260.0 280.0 300.0

RANGE TIME, SECONDS

Figure 9-8. Amplitude and Frequency of Longitudinal Osci 11 ations During S-II Stage Boost

0.30 / 0.20 / AS 505

0.I0 _J Z _FZ'L o 0

I-- -o.lo f __1

-0.20

-0.30

N 22 AS-506 >2 2O AND AS-505 ..__.....--.-----

r_ 16 580 590 600 610 620 630 640 650 RANGE TIME, SECONDS

Figure 9-9. S-IVB AS -506 and AS-505 17- to 20-Hertz Osci I I ati ons Comparison

9-9 ACCELERATION,GIMBALBLOCK(A0012-403) 0.I o.ot,lm ,,M,,q -0.I

=_ ACCELERATION MAIN LOX LINE UPPER END (A0015-424) l,n 0.20 J L_J ¢_)

r_,r, -0.20 580 590 600 610 620 630 640 650

RANGE TIME, SECONDS

Figure 9-10. AS-506 S-IVB First Burn Maximum Response

S-IVB ENGINE GIMBAL BLOCK (A12-403) 0.20 I /_AS-505 FLIGHT TIME #_619 SECONDS

0.15 .....

E 0.10

0.05 TIME

35 FREQUENCY, HERTZ

Figure 9-11. AS-506 and AS-505 First Burn Response

9-I0 During S-IVB second burn, small longitudinal oscillations began on the engine gimbal pad (A12-403) at about 10,164 seconds and peaked (_0.12 g) at 10,172 seconds at the first longitudinal mode frequency of 13 hertz. These oscillations were damped out by 10,184 seconds. Similar 13- to 16-hertz oscillations occurred on AS-505 and other previous flights at approximately the same levels and range time.

The chamber pressure responses were in the noise floor in the 17- to 20- hertz region during first burn and the 13- to 16-hertz region during second burn. The LOX pump inlet and discharge pressure measurements showed in- significantly low amplitudes throughout both S-IVB burns as did the longitudinal accelerometers in the IU and CM.

The data show typical buildup and decay periods of low-level oscillations without indications of propulsion/structural coupling. Since these oscillations have been observed on previous flights, it is assumed that they are characteristic of the stage and could be expected on future flights.

The 45-hertz oscillations that occurred just after the LH2 steD pressurization event on AS-505 were not detected on AS-506. The AS-506 Non Propulsive Vent (NPV) pressures showed very small, ±0.35 N/cm2 (±0.5 psia), pressure oscillations after step pressurization, as shown in Figure 9-12. The IU yaw

AS-506 LH2 NPV PRESSURE (D183-409)

L_ -2 0 AS-506 IU ACCELERATION, YAW (A007-603) 0.2 RANDOM NOISE (TYPICAL TO ECO) 0 -'-vv

-0.2 10,151.0 10,151.5 10,152.0 10,152.5 10,153.0

AS-505 LH2 NPV PRESSURE (D183-409) 8 6 I0 "_

LU 4 LJ 2 C_ C_ 0 0 _- AS-505 IU ACCELERATION_YAW (A007-603) 0.2 45-HERTZ OSCILLATION

-0.2 ....' ' ...... [ .... F"] .... V _"' 9481.0 9481.5 9482.0 9482.5 9483.0 RANGE TIME, SECONDS

Figure 9-12. Comparison of 45-Hertz Oscillations During AS-505 and AS-506 Second Burn

9-11 acceleration measurement (A7-603) showed no response to these small pressure oscillations. In sharp contrast to this condition were the relatively large, _+1.4 N/cm2 (_+2psia), NPVpressure oscillations observed on the AS-505 flight and the resulting 45-hertz vibration indicated by the A7-603 measurement. Therefore, it is assumed that the 45-hertz vibration did not occur on the AS-506 flight.

9.2.3.2 Lateral Dynamic Characteristics. Oscillations in the first four modes were detectable throughout S-IC powered flight. Spectral analyses were performed to determine modal frequencies using 5-second time slices. The frequencies of these oscillations agreed well with the analytical predictions, as shown in Figure 9-13.

LEGEND

ANALYSIS MODE 1 ANALYSIS MODE 2 I0 i i COMMANDMODULE ANALYSIS MODE 3 _'_ I I ANALYSIS MODE4 MEASURED MODE(_) MODE NUMBER N

/ /

N -r- j'" 6 • J' >- /" Z J L_

W ,..J ,,r_ 4 Jd o

--@ F-@--- C_

0 2o 40 60 80 I00 120 140 160 RANGE TIME, SECONDS

Figure 9-13. AS-506 Lateral Analysis/Measured Modal Frequency Correl ati on

9-12 SECTIONI0

GUIDANCEANDNAVIGATION

I0.I SUMMARY

I0.I.I Flight Program

The guidance and navigation system performed satisfactorily at all times for which data are presently available. The parking orbit and translunar injection parameters are well within tolerance. The S-IVB LOX dump, LH2 vent, and Auxiliary Propulsion System (APS) ullage burn resulted in a heliocentric orbit of the S-IVB/IU as planned.

The actual S-II Engine Mixture Ratio (EMR) shift occurred approximately 9.5 seconds later than indicated by the final stage propulsion prediction. About 4 seconds of this deviation was attributed to the change in Launch Vehicle Digital Computer (LVDC) nominal characteristic velocity presetting predictions and variation in actual from predicted flight performance. Approximately 5.5 seconds of the deviation are attributed to improper scaling in the flight program calculation of characteristic velocity. 10.1.2 Instrument Unit Components

The LVDC, the Launch Vehicle Data Adapter (LVDA), and the ST-124M-3 inertial platform functioned satisfactorily. The platform-measured crossrange velocity (Y) exhibited a negative shift of approximately 1.8 m/s (5.9 ft/s) at 3.3 seconds after liftoff. The probable cause was the Y accelerometer head momentarily contacting an internal mechanical stop. Although this had negligible effect on the launch veh'icle, investigation to determine the cause of the velocity shift is continuing. 10.2 GUIDANCECOMPARISONS

The postflight guidance error analysis was based on comparisons of the ST-124M-3 platform measured velocities with the postflight trajectory established from external tracking data. No precision tracking data were available for trajectory construction; therefore, hardware error analysis was limited to gross errors. The comparisons made and reported herein are referenced to the AS-506 final (14 day) postflight trajectory. The boost- to-parking orbit portion of the trajectory was a composite fit of C-Band radar data. The second burn trajectory consists of ST-124M-3 measured velocities constrained to orbital solutions.

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Comparisons of navigation (PACSS13 coordinate system) positions, velocities, and flight path angle at significant flight event times are presented in Table 10-2. The guidance (LVDC) and observed postflight trajectory values are in relatively good agreement for the boost-to-parking orbit burn mode. The initial error in crossrange velocity is reflected in the displacement. The component differences at Time Base 6 and at TLI are still under investigation. There appears to be a timing error of about 2.7 seconds between the orbital solutions and the LVDC. The component differences at parking orbit insertion, Time Base 6, and TLI are given in Table 10-3.

The ST-124M-3 platform measurements and the LVDCflight program were highly successful in guiding the AS-506 vehicle to near nominal end conditions. A minimum of corrections were required for the spacecraft to accomplish a near perfect mission.

10.2.1 Late S-II Stage EMRShift

The S-II stage actual EMRshift occurred approximately 9.5 seconds later than indicated by the final stage propulsion prediction. About 4 seconds of this deviation was attributed to the change in IU LVDCnominal characteristic velocity presetting predictions and variation in actual from predicted flight performance.

About 5.5 seconds of the late EMRshift deviation was due to improper LVDCscaling. The EMRroutine is entered when a time-to-go quantity Tl_, becomes zero or negative. The Tli was larger than predicted because calculated characteristic velocity, upon which Tli is based, was smaller than predicted.

The S-II characteristic velocity calculated in the IU was incorrect due to an unfortunate combination of scaling situations in the LVDCflight program. These scaling errors caused the calculated S-II stage characteristic velocity (BNVC) to be low by approximately 91.7 m/s (300.9 ft/s) at the time EMRshift should have occurred. BNVCis calculated as follows: DVCi2 : (AX2 + Ay2 + AZ2) ABNVC= I/2(DVCi_ 1 + DVCi/DVCi_l) BNVCi = BNVCi_1 + ABNVC

10-4 Table I0-I. Inertial Platform Velocity Comparisons

DATA VEt OClTY M/S (FT/S)** EVENTS SOURCE ALTITUDE (Xm) CROSSRANGE(_m) DOWNRA[IGE (_m)

Guidance 2585.13 1.50 2254.84 (8481.40) (4.92) (7397.77) S-IC Postfl ight 2583.28 2.71 2255.10 OECO Trajectory (8475.33) (8.89) (7398.62) 2239.23 Operational 2600.23 -5.34 Trajectory (8530.94) (-17.52) (7346.56)

Guidance 3430.03 -3.50 6759.48 (11,253.38) (-II.48) (22,176.77) 6759.88 S-ll Postflight 3430.60 -0.96 OECO Trajectory (11,255.25) (-3.15) (22,178.08) 6784.83 Operational 3432.73 -I.I0 Trajectory (11,262.24) (-3.61) (22,259.94)

Guidance 3190.55 1.50 7607.13 (10,467.68) (4.92) (24,957.78) 7607.67 First S-IVB Postfl ight 3192.07 3.23 (24,959.55) ECO Trajectory (10,472.67) (10.60) 7606.29 Operational 3187.05 1.17 Trajectory (10,456.20) (3.84) (24,955.02) 7608.85 Gui dance 3189.85 1.50 (10,465.39) (4.92) (24,963.42) 3191.42 3.29 7609.49 Parking Orbit Postfl ight (24,965.52) Insertion Trajectory (10,470.54) (10.79) 1.18 7607.85 Operational 3186.42 (24,960.14) Trajectory (10,454.13) (3.87)

Guidance 2677.57 275.86 1656.05 (8784.68) (905.05) (5433.23) 1655.22 Second Pos tflight 2677.21 274.0O (5430.51) S-IVB ECO* Trajectory (8783.50) (898.95) 1655.18 Operational 2678.72 276.23 Trajectory (8788.45) (906.27) (5430.38) 1658.50 Guidance 268O.70 276.50 (8794.95) (907.15) (5441.27) 274.79 1658.28 Trans I una r Postflight 2680.38 (5440.25) Injection Trajectory (8793.90) (901.54) 1657.45 Operational 2681.45 276.77 Trajectory (8797.41) (908.04) (5437.83)

| *Second burn velocity data represent accumulated velocities from Time Base 6. **PACSS 12 Coordinate System.

10-5 Table 10-2. Guidance Comparisons

POSITIONS VELOCITIES FLIGHT PATH (METERS) M/S DATA ANGLE (BEG) EVENT (FT) (FT/S) SOURCE

Xs Vs Zs R Xs Ys is Vs "i

Guidance 6,437,364 39,169 159,182 6,439,450 841.25 121.94 2630.38 2764.32 19.1484 (21,119,961) (128,507) (522,251 (21,126,804) (2760.01) (400.07) (8629.86) (9069.29) S-IC Postflight 6,437,247 39,441 159,164 6,439,335 839.62 123.11 2630.64 2764.13 19.1143 OECO Trajectory (21,119,577) (129,400) (522,192) (21,126,426) (2754.66) (403.90) (8630.71) (9068.67) Operational 6,437,901 38,968 157,58( 6,439,948 861.66 115.11 2614.93 2755.64 19.635 Trajectory (21,121,724) (127,847) (517,0141 (21,128,437) (2826.97) (377.66) (8579.17) (9040.81)

Guidance 6,289,965 79,413 1,860,94_ 6,559,961 -1891.78 88.17 6651.61 6915.96 0.6139 (20,636,368) (260,541) (6,105,471 (21,522,182) (-6206.63) (289.27) (21,822.87) (22,690.16.)

Postflight 6,289,873 80,367 1,859,72( 6,559,537 -1891.43 90.43 6651.87 6916.14 0.6075 OECO Trajectory (20,636,067) (263,670) (6,101,444] (21,520,790) (-6205.48) (296.69) (21,823.72) (22,690.75) I S-II -1917.77 1 Operational 6,283,160 79,489 1,884,67! 6,560,214 90.69 6668.07 6938.96 0.661 Trajectory (20,614,043) (260,791) (6,183,310](21,523,012) (-6291.90) (297.54) (21,876.87) (22,765.62) o -3433.60 ! Guidance 5,891,469 91,789 2,891,28! 6,563,335 i 76.86 6993.63 7791.43 -0.00148 (-11,265.09) (19,328,968) (301,144) (9,485,842](21,533,251) i (252.17) (22,944.98) (25,562.43) First Postflight 5,890,834 93,058 2,892,01 6,563,105 -3432.48 78.03 6993.87 7791.17 0.01511 S-IVB ECO Trajectory (19,326,885) (305,310) (9,488,2451 (21,532,498) (-11,261.42) (256.00) (22,945.77) (25,561.58) Operational 5,890,252 91,860 6,563,311 -3436.57 76.94 6992.15 7791.41 -0.00220 Trajectory (19,324,973) (301,377) (21,533,172) (-11,274.84) (252.43) (22,940.12) (25,562.39)

Guidance 5,856,709 92,552 6,563,334 -3517.14 75.69 6954.08 7793.28 -0.00064 (19,214,926) (303,647) (21,533,247) (-II,539.17) (248.33) (22,815.22) (22,568.50)i

Parking Orbit Postflight 5,856,252 93,832 2,961,27( 6,563,052 -3515.97 76.90 6954.42 7793.07 0.01205 Insertion Trajectory (19,213,427) (307,846) (9,715,472)(21 ,532,323)I(-II,535.33) (252.30) (22,816.34) (22,567.81)

Operational 5,855,466 92,623 2,963,43E 6,963,309 -3520.02 75.77 6952.42 7793.10 -0.00142 Trajector_ (19,210,845) (303,882) (9,722,567)(21 ,533,166) (-II,548.62) (248.59) (22,809.78) (22,567.91)

Guidance -2,290,189 -142,062 6,570,377 7301.51 -26.14 -2720.93 7792.02 0.03779 (-7,513,744) (-466,083) 21,556,355) (23,955.09) (-85.76) (-8926.94) (25,564.50)

Time Base 6 Postflight -2,313,353 -143,619 -6,149,910 6,572,186 7290.84 -26.54 -2745.81 7790.80 0.02690 Trajectory (-7,589,739) (-471,191) -20,176,871 ) 21,562,288) (23,920.08) (-87.07) (-9008.56) (25,560.37) Operational -2,322,832 -142,124 -6,148,429 6,574,110 7284.60 -26.96 -2756.71 7788.81 O. 03622 Trajectory (-7,620,838) (-466,286) -20,172,011) 21,568,603) (23,899.61) (-88.45) (-9044.32) (25,553.84) Table I0-2. Guidance Comparisons (Continued)

POSITIONS VELOCITIES FLIGHT PATH (METERS) M/S ANGLE (DEG) DATA (FT) (FTIS) EVENT SOURCE Ys Zs V s Xs Ys Zs Xs

6.98785 6,699,493 8393.63 408.79 6847.05 10,839.84 Guidance 4,836,7441 -61,604 -4,635,225 (1341.17) (22,464.07) (35,563.78) (-202,I12) -15,207,431) (21,979,963): (27,538.16) (15,868,583) 8412.93 408.40 6825.11 10,840.96 6.91287 Second Postflight 4,817,420' -63,231 -4,653,78E 6,698,451 (27,601.48) (1339.90) (22,392.09; (35,567.45) S-IVB ECO Trajectory (15,805,184) (-207,451) -15,268,322) 121,976,545) 6825.36 10,838.72 6.959 -4,651,489 6,701,348 8409.82 408.93 Operational 4,823,682 -61,683 (21,986,049) (27,591.27) (1341.63) (22,392.91) (35,560.]0) Trajectory (15,825,729) (-202,371) (-15,260,971) 7.44149 8332.21 410.24 6910.34_ 10,832.68 Gui dance 4,920,373 -57,507 6,713,101 (27,336.651 (1345.93) (22,671.72) (35,540.29) (16,142,956) (-188,670) (22,024,610) 6889.31 10,834.31 7.36695 6,711,964 8351.74 410.04 Translunar Postflight 4,901,502 -59,131 (22,020,878) (27,400.72) (1345.28) (22,602.72) (35,545 64)! Injection Trajectory (16,081,042) (-193,998) iJ 10,831.12 7.412 o 6,714,891 8348.15 410.29 6888.63 Operational 4,907,485 -57,584 -4,582,90; I (22,030,484) (27,388.94) (1346.10) (22,600.49) (35,535.17) Trajectory 16,100,674) (-188,924) -15,035,785: Table 10-3. Guidance Components Differences

PARAMETERS OBSERVED TRAJECTORY OPERATIONAL TRAJECTORY LVDC LVDC

PARKING ORBIT INSERTION

A_s m/s (ft/s) 1.17 (3.84) -4.05 ('-13.29) A_>s m/s (ft/s) 1.21 (3.97) 0.08 (0.26) A_s m/s (ft/s) 0.34 (I.12) -1.66 (-5.45) AVs m/s (ft/s) -0.21 (-0.69) -0.18 (-0.59) AXs m (ft) -457 (-1499) -1244 (-4081) Ay m (ft) 1280 (4199) 72 (236) AZ m (ft) 239 (784) 2402 (7880) AR m (ft) -282 (-925) -25 (-82) A@ deg 0.01269 -0.00098

TIME BASE 6

AXs m/s (ft/s) -10.67 (-35.01) -16.91 (-55.48) A#s m/s (ft/s) -o.4o (-i.31) -0.82 (-2.69) AZs m/s (ft/s) -24.88 (-81.63) -35.78 (-117.39) AVs m/s (ft/s) -1.26 (-4.13) -3.25 (-I0.66) Ax s m (ft) -23,163 (-75,994) -32,642 (-I07,093) Ay s m (ft) -1557 (-5108) -62 (-203) AZs m (ft) 6770 (22,210) 8251 (27,070) ARm (ft) 1808 (5932) 3733 (12,247) AO de9 -0.01089 -0.00157

TRANSLUNAR INJECTION

AAs m/s (ft/s) 19.53 (64.07) 15.94 (52.30) A#s m/s (ft/s) -0.20 (-0.66) 0.05 (0.16) AZs m/s (ft/s) -21.03 (-69.00) -2| .71 (-71.23) AVs m/s (ft/s) 1.63 (5.35) -1.56 (-5.12) AXs m (ft) -18,871 (-61,913) -12,888 (-42,283) Ay s m (ft) -1624 (-5328) -78 (-256) AZs m (ft) -18,564 (-60,906) -16,470 (-54,035) ARm (ft) -1138 (-3734) 1790 (5873) A@ deg -O.O7454 -0.03548

10-8 During the calculation of BNVC, DVCi2 lost significant information due to round off. The scaling factor required 28 bits of information. The LVDC maintains only 25 bits of information. The three least significant bits of information were lost by computer underflow. Another bit was lost due to the binary arithmetic and hardware algorithm for division. As a consequence, the apparent increase in BNVC per computer cycle was less than the actual increase in BNVC. The total S-II characteristic velocity error at the time of actual EMR shift was about 92.6 m/s (303.8 ft/s). The deviation in EMR shift time caused no performance perturbation. Figure 10-3 gives the differences between S-II stage correct BNVC values and those computed by the LVDC flight program. Investigation is being conducted to improve scaling in IU LVDC velocity calculations.

10.3 NAVIGATION AND GUIDANCE SCHEME EVALUATION

All analyzed guidance and navigation measurements indicated satisfactory performance throughout the flight. The active guidance phases start and stop times are given in Table 10-4. Included in this table are the start and stop times in the artificial tau phases and chi freezes. The rate- limited attitude commands shown in Figures 10-4 and 10-5 indicate near nominal performances.

The flight program routine causing S-II EMR shift to be commanded was entered later than predicted in the OT. This deviation is discussed in paragraph 10.2.1.

120 390

360

330 100 / ._ _ _ / 300 __

/ Ir 270 _= 80 /f ._/ 240 /_f 210 ,,J , / , 60 / 180

150

40 120 o o O, 90

20 60

0 / 30 1O0 150 200 250 300 350 400 450 500 550 RANGE TIME, SECONDS

Figure 10-3. AS-506 Characteristic Velocity Error

10-9 Table 10-4. Start and Stop Times for IGM Guidance Commands

STEERING TERMINAL IGM PHASE ARTIFICIAL TAU MISALIGNMENT CHI FREEZE EVENT* GUIDANCE CORRECTION (SEC) (SEC) (SEC) (SEC) (SEC)

Start Stop Start Stop Start Stop Start Stop Stop Start

First Phase IGM 204.07 494.83 221.52 496.82

Second Phase IGM 494.83 548.2 494.83 504.16 498.73 548.2

Third Phase IG!_ 548.2 691.64 555.57 562.43 562.43 691.64 665.15 691.64 691.64 699.26**

Fourth Phase IG_a 9862.51 9974.58 9_72.87

Fifth Phase IGM 9974.58 10,201.92 9974.58 9980.37 10,199.70 10,174.47 10,201.92 I0,201.92 I0,203.56"*

* All times are for the start of the computation cycle in which the event occurred,

** Start orbital timeline.

Orbital guidance events were accomplished satisfactorily. All S-IVB stage first and second burn guidance parameters indicate satisfactory operation. The orbital insertion conditions after S-IVB first burn are given in Table 10-5, The TLI parameters after S-IVB second burn are given in Table 10-6.

10.4 GUIDANCE SYSTEM COMPONENT EVALUATION

10.4.1 LVDC Performance

The LVDC performed as predicted for the AS-506 mission. No valid error monitor word and no self-test error data have been observed that indicate any deviation from correct operation,

10.4.2 LVDA Performance

The LVDA performance was nominal. No valid error monitor words and no self-test error data indicating deviations from correct performance were observed.

10.4.3 Ladder Outputs

The ladder networks and converter amplifiers performed satisfactorily. No data have been observed that indicate an out-of-tolerance condition between Channel A and the reference channel converter-amplifiers.

10.4.4 Telemetry Outputs

Analysis of the available LVDA telemetry buffer and flight control computer attitude error plots indicated symmetry between the buffer outputs and the ladder outputs. The available LVDC power supply plots indicates satisfactory power supply performance, The H60-603 guidance computer telemetry was completely satisfactory.

I0-I0 _7 S-IC CECO _7 S-II EMR SHIFT _7 S-IC/S-II SEPARATION _7 S-II/S-IVB SEPARATION _7 IGM INITIATION _7 BEGIN CHI BAR STEERING _7S-IC CECO _7S-II EMR SHIFT _7 S-II CECO S_7 GUIDANCE CUTOFF _TS-IC/S-II SEPARATION _TS-II/S-IVB SEPARATION _71GM INITIATION _7BEGIN CHI BAR STEERING _OPERATIONAL TRAJECTORY ..... LVDC 1.4 _TS-II CECO S_TGUIDANCE CUTOFF tA_ _OPERATIONAL TRAJECTORY ---LVDC I0 1.2 I I 0 I I 1.0 I -I0 I I I -20 I 0.8

-30

-40 0.6

c_ o -50 # l o.4 / I -6O 8 ! S / ( -70 I 0.2 / / p- / -80 \ / / / / -90 / / ,! /" -I00 \ / / / -0.2 -II0 I l -120 _7_7 _' _' -o.4 W ' '7 '_' _ _' -I00 0 1O0 200 300 400 500 600 700 800 -I O0 l O0 200 300 400 500 600 700 8OO

RANGE TIME, SECONDS RANGE TIME, SECONDS

Figure 10-4. Attitude Commands During Active Guidance Period EVENT SEQUENCE I0_ x_7 IGM INITIATION EMR SHIFT BEGIN CHI BAR STEERING

BEGIN CHI FREEZE I_ 8_ OPERATIONAL TRAJECTORY /'/ "I --- LVDC T

'I I I /f 6 I d I

o o 4 I /_

l,i r_ , .9 / F- ,! ._ V

I 0 ii, V"

-2

i -16 i

-2O

-_ -24 \ i.- -28

0 _ -32

-36

-4O I

-44 9840 9880 9920 9960 I0,000 10,040 10,080 10,120 I0,160 10,200

RANGE TIME, SECONDS iR7, , ,_z , , , , ,_R7 2:44:00 2:44:40 2:45:20 2:46:00 2:46:40 2:47:.20 2:48:00 2:48:40 2:49:20 2:50:00

RANGE TIME, HOURS: MINUTES : SECONDS

Figure 10-5. Attitude Angles During S-IVB Second Burn

I0-12 Table 10-5. Parking Orbit Insertion Parameters

OPERATIONAL POSTFLIGHT TRAJECTORY LVDC PARAMETER TRAJECTORY TRAJECTORY MINUS OT LVDC MINUS OT

Inertial Velocity 7793.3 m/s 7793.1 7793.1 0.0 0.18 (25,568.5) (ft/s) (25,567.9) (25,567.9) (0.0) (0.6)

Flight Path Angle -0.0006 0.0004 deg -0.001 0.012 0.013

Descending Node -0.012 123.102 0.002 deg 123.100 123.088

Inclination -0.010 32.532 O.OOl deg 32.531 32.521 -0. 00001 0.00007 -0.00015 Eccentricity 0.00022 0.00021

10.4.5 Discrete Outputs

No valid discrete output register words (tags 043 and 052) were observed to indicate guidance or simultaneous memory failure.

10.4.6 Switch Selector Functions

Switch selector data indicate that the LVDA switch selector functions were performed satisfactorily. No error monitor words were observed that indicate disagreement in the Triple Modular Redundant (TMR) switch selector register positions or in the switch selector feedback circuits. No mode code 24 words or switch selector feedback words were observed that indicated a switch selector feedback was in error. In addition, no indications were observed to suggest that the B channel input gates to the switch selector register positions were selected.

Table 10-6. Translunar Injection Parameters

OPERATIONAL POSTFLIGIIT TRAJECTORY LVDC PARAMETER TRAJECTORY TRAJECTORY NINUS OT LVDC MINUS OT

Inertial Velocity m/s 10,831.1 10,834.3 3.2 I0,832.7 1.6 (ft/s) (35,535.1) (35,545.6) (10.5) (35,540.4) (5.3)

Descending Node deg 121.866 121.847 -0.019 121.855 -0.011

Inclination deg 31.379 31.383 0.004 31.382 0.003

Eccentri city 0.97667 0.97696 0.00029 0.97670 0.00003

c3 m2/s2 -1,408,484 -1,391,607 16,877 -1,406,545 1939 (ft2/s 2 ) (-15,160,796) (-14,979,133) (181,633) (-15,139,924) (20,872)

10-13 10.4.7 ST-124M-3 Inertial Platform

The inertial platform system performed as designed. At 3.3 seconds after liftoff, the Y velocity (crossrange) exhibited a change of approximately -1.8 m/s (-5.9 ft/s). A lack of data prevents a precise determination of the cause; however, the probable cause was the Y accelerometer head momentarily contacting an internal mechanical stop. The forcing function was probably crossrange polarized 35 to 40 hertz vibrations, the natural frequency of the accelerometer servo loop.

The Y accelerometer head movement indicated significant incident vibrations. However, the measurement was sampled rather than continuous so the frequency and amplitude of the head motion cannot be readily defined.

The Y gyro was relatively unperturbed, but the X and Z gyros showed significant activity. This indicates a forcing function, probably vibration, mainly along the platform Y axis.

The body-mounted yaw accelerometer, A7-603, was oriented in the same direction as the Y accelerometer. It indicated a generally high random level of vibration which included significant amplitudes between 35 and 45 hertz. The amplitude is presently indeterminate because of telemetry channel and band width limitations.

The X, Y, and Z gyro servo loops for the stable element functioned as designed. The operational limits of the servo loops were not reached at anytime during the mission.

The inertial gimbal temperatures fell below specifications; however, there are no indications of degraded inertial performance. The temperature went below the minimum specification of 313.15°K (104.0°F) at 15,600 seconds, reaching 312.2°K (I02.3°F) at approximately 20,800 seconds.

10-14 SECTIONII

CONTROLSYSTEM

II.I SUMMARY

The AS-506 control system, which was essentially the same as that of AS-505, performed satisfactorily. The Flight Control Computer (FCC), Thrust Vector Control (TVC), and Auxiliary Propulsion System (APS) satisfied all requirements for vehicle attitude control during the flight. Bending and slosh dynamics were adequately stabilized. The prelaunch programed yaw, roll, and pitch maneuvers were properly executed during S-IC boost.

During the maximumdynamic pressure region of flight, the launch vehicle experienced winds that were less than 95-percentile July winds. The maximumaverage pitch and yaw engine deflections were the result of wind shears.

S-IC/S-II first and second plane separations were accomplished with no significant attitude deviations. At Iterative Guidance Mode (IGM) initiation, a pitch up transient occurred similar to that seen on previous flights. At S-II early Center Engine Cutoff (CECO), the guidance parameters were modified by the loss of thrust. There was a change in yaw attitude due to the slight thrust misalignment of the center engine. S-II/ S-IVB separation occurred as expected and without producing any significant attitude deviations.

Satisfactory control of the vehicle was maintained during first and _econd S-IVB burns and during coast in Earth Parking Orbit (EPO). During the Commandand Service Module (CSM) separation from the S-IVB/IU and during the Transposition, Docking and Ejection (TD&E) maneuver, the control system maintained the vehicle in a fixed inertial attitude to provide a stable docking platform. Following TD&E, S-IVB/IU attitude control was maintained during the maneuver to the slingshot attitude and during the LOX dump and LH2 vent. 11.2 S-IC CONTROLSYSTEMEVALUATION

The AS-506 control system performed adequately during S-IC powered flight. The vehicle flew through winds which were less than 95 percentile for July in the maximumdynamic pressure region of flight. Less than I0 percent of the available engine deflection was used throughout flight (based on average engine gimbal angle). S-IC outboard engine cant was accomplished as planned.

II-I All dynamics were within vehicle capability. In the region of high dynamic pressure, the maximumangles-of-attack were 1.6 degrees in pitch and 1.4 degrees in yaw. The maximumaverage pitch and yaw engine deflections were 0.2 degree and 0.3 degree, respectively, in the maximumdynamic pressure region. Both deflections were due to wind shears. The absence of any divergent bending or slosh frequencies in vehicle motion indicates that bending and slosh dynamics were adequately stabilized. Vehicle attitude errors required to trim out the effects of thrust imbalance, thrust misalignment, and control system misalignments were within predicted envelopes. Vehicle dynamics prior to S-IC/S-II first-plane separation were within staging requirements. 11.2.1 Liftoff Clearances

The launch vehicle cleared the mobile launcher structure within the available clearance envelopes. Camera data showing liftoff motion were not available for the AS-506 flight, but simulations with flight data show that less than 15 percent of the available clearance was used. The ground wind was from the south with a magnitude of 3.3 m/s (6.4 knots) at the 18.3 m (60 ft) level.

The predicted and measured misalignments, slow release forces, winds, and the thrust-to-weight ratio are shown in Table II-I.

11.2.2 S-IC Flight Dynamics

Maximumcontrol parameters during S-IC burn are listed in Table 11-2. Pitch, yaw, and roll plane time histories during S-IC boost are shown in Figures II-I, 11-2, and 11-3. Dynamics in the region between liftoff and 40 seconds result primarily from guidance commands. Between 40 and II0 seconds, maximumdynamics were caused by the pitch tilt program, wind magnitude, and wind shears. Dynamics from II0 seconds to S-IC/S-II separation were caused by high altitude winds, separated air flow aerodynamics, center engine shutdown, and tilt arrest. The transient at CECOindicates that the center engine cant was 0.2 degree in yaw and -0.06 degree in pitch.

At Outboard Engine Cutoff (OECO), the vehicle had attitude errors of -0.3, 0.I, and 0.0 degree in pitch, yaw, and roll, respectively. These errors are required to trim out the effects of thrust imbalance, offset Center of Gravity (CG), thrust vector misalignment, and control system misalignments. The maximumequivalent thrust misalignments were 0.II, -0.05 and -0.02 degree in pitch, yaw, and roll, respectively.

There was no significant sloshing observed. The engine response to the observed slosh frequencies showed that the slosh was well within the capabilities of the control system.

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PREFLIGHT PREDICTED LAUNCH

PITCH YAW ROLL PITCH YAW ROLL Thrust Misalignment ±0.34 +0.34 ±0.34 O.ll -0.05 -0.02 deg* Center Engine -0.06 0.2 Cant, deg Servo Amplifier +0.I ±0.I +0.1 Offset, deg/eng Vehicle Stacking & ±0.29 +0.29 0.0 0.12 -0.06 0.0 Pad Misalignment, deg Attitude Error at 0.06 -0.02 0.02 Holddown Arm Release, deg

Peak Slow Release 415,000 (93,300) 400,000 (90,000) *** Force Per Rod, N (Ibf)

Wind 95 Percentile Envelope 3.3 m/s (6.4 Knots) At 18.3 Meters (60 Feet) Thrust to Weight 1.195 Ratio

*Thrust misalignment of 0.34 degree encompasses the center engine cant. A positive polarity was used to dete_line minimum fin tip/umbilical tower clearance. A negative polarity was used to determine vehicle/GSE clearances. **Data not available for update. *** Approximate data obtained during a data dropout period.

The normal accelerations observed during S-IC burn are shown in Figure 11-4. Pitch and yaw plane wind velocities and angles-of-attack are shown in Figure 11-5. The winds are shown both as determined from balloon and rocket measurements and as derived from the vehicle Q-ball.

11.3 S-II CONTROL SYSTEM EVALUATION

The S-II stage attitude control system performance was satisfactory. Analysis of the magnitude of modal components in the engine deflections revealed that vehicle structural bending and propellant sloshing had negligible effect on control system performance. The maximum values of pitch and yaw control parameters occurred in response to IGM Phase 1 initiation. The maximum values of roll control parameters occurred in

11-6 Table 11-2. Maximum Control Parameters During S-IC Burn

PITCH PLANE YAW PLANE ROLL PLANE

RANGE RANGE RANGE PARAMETERS UNITS MAGNITUDE TIME MAGNITUDE TIME MAGNITUDE TIME (SEC) (SEC) (SEC)

Attitude Error deg 0.83 117.4 -I .02 3.3 -0.92 13.8

Angular Rate deg/s -0.97 69.1 -0.53 12.6 1.38 14.4

Average Gimbal deg 0.23 90.6 -0.44 3.2 -0.09 80.9 Angle

Angle-of-Attack de9 1.82 117,1 1.42 73.0

Angle-of-Attack/ deg-N/cm 2 4.93 91.4 4.50 73.0 Dynamic Pressure Product

Normal m/s 2 -0.331 95.5 0.306 63.9 - Acceleration

response to S-IC/S-II separation disturbances. The control responses at other times were within expectations.

Between the events of S-IC OECO and initiation of IGM, the vehicle attitude commands were held constant. Significant events occurring during this interval were S-IC/S-II separation, S-II stage J-2 engine start, second plane separation, and Launch Escape Tower (LET) jettison. The attitude control dynamics throughout this interval indicated stable operation as shown in Figures 11-6 through 11-8. Steady-state attitudes were achieved within 20 seconds from S-IC/S-II separation. The maximum control parameter values for the period of S-II burn are shown in Table 11-3.

At IGM initiation, the TVC received FCC commands to pitch the vehicle up. During IGM, the vehicle pitched down at a constant commanded rate of approximately -0.I deg/s. The transient magnitudes experienced at IGM initiation were similar to those of the AS-504 and AS-505 flights.

A steady-state pitch attitude error of approximately 0.15 degree resulted from thrust imbalance. Following CECO, a steady-state yaw attitude error of -0.3 degree occurred. Pea_ transient yaw attitude error after CECO was -0.5 degree at 464 seconds. This yaw error occurred in response to the loss of the compliance deflection of the center engine at cutoff. The center engine was not precanted to allow for compliance deflection. This compliance effect occurred in the yaw plane because of the location of the fixed links. Consequently, the outboard engines were deflected in yaw after CECO to compensate for the yaw attitude error and to stabilize the vehicle. The deflections of the outboard engines in pitch after CECO were the result of a pitch-up guidance command. This command was generated to compensate for the effects that loss of center engine thrust would have upon the flight trajectory.

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II-II AVERAGEROLLENGINE ROLLBODYRATE ROLLATTITUDEERROR DEFLECTION(POSITIVECCW (POSITIVECWVIEWED (POSITIVECWVIEWED VIEWEDFROMREAR),deg FROMREAR),deg/s FROMREAR),deg

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PITCH PLANE YAW PLANE ROLL PLANE

RANGE RANGE RANGE PARAMETERS UNITS MAGNITUDE TIME MAGNITUDE TIME MAGNITUDE TIME (SEC) (SEC) (SEC)

Attitude Error deg -I.9 206.9 -0.54 464.8 -1.6 165.5

Angular Rate deg/s 1.2 207.8 0.2 467.0 1.7 166.2

Average Gimbal deg -0.9 165.3 -0.33 464.0 -0.54 165.3 Angle

II.4 S-IVB CONTROL SYSTEM EVALUATION

The S-IVB TVC system provided satisfactory pitch and yaw control during powered flight. The APS provided satisfactory roll control during first and second burns.

During S-IVB first and second burns, control system transients were experienced at S-II/S-IVB separation, guidance initiation, Engine Mixture Ratio (EMR) shift, chi bar guidance mode, and J-2 engine cutoff. These transients were expected and were within the capabilities of the control system.

11.4.1 Control System Evaluation During First Burn

The S-IVB first burn attitude control system response to guidance commands for pitch, yaw and roll are presented in Figures 11-9, II-I0 and II-II, respectively. The maximum attitude errors and rates occurred at IGM and chi bar steering initiation. A summary of maximum values of the critical flight control parameters during S-IVB first burn is presented in Table 11-4.

The pitch and yaw effective thrust vector misalignments during first burn were 0.22 and -0.33 degree, respectively. A steady-state roll torque of 61.4 N-m (45.3 Ibf-ft), counterclockwise looking forward, required roll APS firings during first burn. The steady-state roll torque experienced on previous flights has ranged between 27 N-m (20 Ibf-ft) counterclockwise and 54.2 N-m (40.0 Ibf-ft) clockwise.

II.4.2 Control System Evaluation During Parking Orbit

The coast attitude control system provided satisfactory orientation and stabilization of the S-IVB/CSM in parking orbit. The only maneuver during parking orbit was to align the vehicle with the local horizontal just after S-IVB first cutoff. Pitch axis control parameters during the maneuver to the local horizontal are indicated in Figure 11-12. The yaw and roll control parameters did not show significant transients and are not presented.

11-13 _7S-IVB ESC S_71GM PHASE 3 INITIATED S_7BEGIN CHI BAR STEERING _7CHI FREEZE INITIATED S_TS-IVB FIRST CUTOFF ,., 2 O0

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11.4.3 Control System Evaluation During Second Burn

The S-IVB second burn attitude control system response to guidance commands for pitch, yaw and roll are presented in Figure 11-13, 11-14 and 11-15, respectively. The maximum attitude errors and rates occurred at guidance initiation and EMR shift. A summary of maximum values of the critical flight control parameters during S-IVB second burn is presented in Table 11-5.

The pitch and yaw effective thrust vector misalignments during second burn were approximately 0.25 and -0.35 degree, respectively. The steady-state roll torque during second burn ranged from 42.1 N-m (31.1 Ibf-ft) at the low EMR to 52.3 N-m (38.6 Ibf-ft) at the 5.0:1.0 EMR.

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PITCH PLANE YAW PLANE ROLL PLANE

RANGE RANGE RANGE PARAMETERS UNITS MAGNITUDE TIME MAGNITUDE TIME MASNITUDE TIME (SEC) (SEC) (SEC)

Attitude Error deg +1.80 668.1 -0.77 562.7 -1.15 584.2

Angular Rate degls +0.90 560.1 -0.29 560.0 +0.I0 554.3

Gimbal Angle deg +1.08 562.7 -0.79 561.5

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11-17 S-IVB EMR SHIFT S-IVB SECOND CUTOFFESC

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0"0L8'6 66"0- £'L98'6 9t'L 6ep eL6uV LEqm#9 i_'L£8'6 SLL'O 0"698'6 O£'O- 9"898'6 £Z'L- s/6ap a%_8 JelnSuv

0"0L6'6 60" L- 0"#86'6 £Z'L- 6"L98'6 E'_ 5ep _o_J3 apn%_%%V

(3as) (335) (33S) 3NI$ 30nlINgVN 3Wll 3ORIlN9VN 3Wll 30nllNgVH SIINN S_313NV_Vd 39NV'd 39NV_ 39NV_

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3anLiLLv 3!_a± o± a3An3NVW3±VILINIZ_ _INITIATE MANEUVER TO SLINGSHOT ATTITUDE 0_ 4.0

hlJ 0.0

I--- p--4 ,-.-q v) • "r" F--O -4.0 o_ z,i 1.0 _..JO

r,,- L.LI 0.0

-1.0

e._. LL1 OV') 4.0

0.0 J I_ _.-.4 I---

I...- 0 (-_ I-- _. _---_ -4.0 1.0

i 0.0

-I .0

0 4.0 1

J_ 0.0 ___- I-- _i) _z ,-,," -4.0 .._.1 l.O

0.0 .-.J

01._ -l .0 17,380 17,420 17,460 17,500 17,540 17,580 RANGE TIME, SECONDS

i I I I I _ I I I I l ! 04 :49 :40 04:51:20 04:58:00 RANGE TIME, HOURS:MINUTES:SECONDS

Figure ll-18. Pitch, Yaw and Roll Plane Dynamics During The Maneuver to Slingshot Attitude

ll-21/ll-22

SECTION12

SEPARATION

12.1 SUMMARY

S-IC/S-II first plane separation was satisfactory. Related data indicate that the S-IC retromotors performed as expected. Similarly, S-II second plane separation and S-II/S-IVB separation were nominal. The S-II retromotors and S-IVB ullage motors performed as expected. Commandand Service Module (CSM) separation from the Launch Vehicle (LV) occurred as predicted during translunar coast. The Transposition, Docking, and Ejection (TD&E) maneuver occurred as expected. Attitude control of the LV was maintained during each separation sequence.

12.2 S-IC/S-II SEPARATIONEVALUATION

S-IC/S-II separation and associated sequencing were accomplished as planned. Dynamic conditions at separation were within staging limits. Rate gyros and accelerometers located on the Instrument Unit (IU) showed no disturbances, indicating a clean severance of the stages. Data from the Exploding Bridge Wire (EBW) firing unit indicate that S-IC retromotor ignition was accomplished. The S-II ullage motors performed as predicted. Since there were no cameras on the S-II stage, calculated dynamics of the interstage and the S-II stage were used to determine if second plane separation was within the staging requirements. 12.3 S-II/S-IVB SEPARATIONEVALUATION

The S-II retromotors and the S-IVB ullage motors performed satisfactorily and provided a nominal S-II/S-IVB separation. Dynamic conditions at separation were within staging, limits with separation conditions similar to those observed on previous flights.

12.4 S-IVB/IU/LM/CSM SEPARATIONEVALUATION

Separation of the CSMfrom the LV was accomplished as planned. There were no large control disturbances noted during the separation.

12-I 12.5 LUNARMODULEDOCKINGAND EJECTIONEVALUATION

The attitude of the LV was adequately maintained during the docking of the CSMwith the Lunar Module (LM). The CSM/LMwas then successfully spring ejected from the LV. There were no significant control disturbances during the ejection.

12-2 SECTION13

ELECTRICAL NETWORKS

13.1 SUMMARY

The AS-506 launch vehicle electrical systems performed satisfactorily throughout all phases of flight. Operation of the batteries, power supplies, inverters, Exploding Bridge Wire (EBW) firing units, switch selectors, and interconnecting cabling was normal. 13.2 S-IC STAGEELECTRICALSYSTEM

The voltage for Battery No. 1 (Operational) and Battery No. 2 (Instrumen- tation) remained within performance limits of 26.5 to 32.0 vdc during powered flight. Battery currents were near predicted and below the maximum limit of 64 amperes for both Battery No. 1 and Battery rlo. 2. Battery power consumption was well within the rated capacity Of 640 ampere-minutes for both Battery No. 1 and Battery No. 2, as shown in Table 13-I. The two measuring power supplies remained within the 5 +_0.05 vdc design limit during powered flight.

Table 13-I. S-IC Stage Battery Power Consumption

POWERCONSUMPTION*

RATED PERCENT BUS CAPACITY OF BATTERY DESIGNATION (AMP-MIN) AMP-MIN CAPACITY

Operational No. 1 IDIO 640 29.6 4.6 Instrumentation No. 2 1D20 640 90.0 14.1

*Battery power consumptions were calculated from power transfer until S-IC/S-II separation.

13-I All switch selector channels functioned correctly, and all outputs were issued within their required time limits in response to commandsfrom the Instrument Unit (IU). The separation and retromotor EBWfiring units were armed and triggered as programed. Charging times and voltages were within tile requirements of 1.5 seconds for maximumallowable charging time and 4.2 +_0.4 volts for the allowable voltage level.

The commanddestruct EBWfiring units were in the required state of readiness if vehicle destruct became necessary. 13.3 S-II STAGEELECTRICALSYSTEM

All battery bus voltages remained within specified limits throughout the prelaunch and flight periods, and bus currents remained within required and predicted limits. Main bus current averaged 36 amperes during S-IC boost and varied from 49 to 57 amperes during S-II boost. Instrumentation bus current averaged 23 amperes during S-IC and S-II boost. Recirculation bus current averaged 97 amperes during S-IC boost, and ignition bus current averaged 31 amperes during the S-II ignition sequence. Battery power consumption was well within the rated capacities of the batteries as shown in Table 13-2.

Table 13-2. S-II Stage Battery Power Consumption

BUS RATED POWERCONSUMPTION* TEMPERATURE DESIG- CAPACITY PERCENTOF BATTERY NATION (AMP-MR) AMP-HR CAPACITY MAX MIN

Main 2DI1 35 7.96 22.7 3O5.4oK 299.8°K (90.O°F) (80.O°F) Ins t rumentati on 2D21 35 3.80 10.9 301.5°K 299.5°K (83.0°F) (79.5°F) Recircul ati on 2D51 3O 5.68 18.9 302.9°K 299.8°K No. 1 (85.5°F) (80.O°F) Recircul ati on 2D51 3O 5.73 19.1 307.6°K 304.3°K No. 2 and (94.0°F) (88.0°F) 2D61

*Power consumption calculated from -50 seconds.

13-2 The five temperature bridge power supplies and the three 5-vdc instrumentation power supplies all performed within acceptable limits. The five LH2 recirculation inverters that furnish power to the recirculation pumps operated properly throughout the J-2 engine chilldown period. All switch selector channels functioned correctly, and all outputs were issued within their required time limits in response to commandsfrom the IU. Performance of the EBWcircuitry for the separation system was satisfactory. Firing units charge and discharge responses were within predicted time and voltage limits. The commandEBWfiring units were in the required state of readiness if vehicle destruct became necessary. 13.4 S-IVB STAGEELECTRICALSYSTEM

The voltages, currents, and temperatures of the three 28-vdc and one 56-vdc batteries stayed well within acceptable limits as shown in Figures 13-I through 13-4. Battery temperatures remained below the 322°K (120°F) limit during the powered portions of flight. (This limit does not apply after insertion into orbit.) The highest temperature of 316.5°K (II0 F) was reached on Aft Battery No. 2, Unit I, after S-IVB first burn cutoff. Battery power consumption is shown in Table 13-3. All switch selector channels functioned correctly, and all outputs were issued within their required time limits in response to commandsfrom the I U.

Table 13-3. S-IVB Stage Battery Power Consumption

RATED POWERCONSUMPTION** CAPACITY PERCENTOF BATTERY (AMP-HRS)* AMP-HRS CAPACITY

Forward No. 1 300.0 121.5 40.5

Forward No. 2 24.75 25,4 102.7

Aft No. 1 300.0 78.2 26.1

Aft No. 2 75.0 42.3 56.4

*Rated capacities are minimum guaranteed by vendor. **Actual usage to 29,000 seconds (08:03:20) is based on flight data.

13-3 (DTRANSFER TO INFE_AL (_FORWARD BATTERY NO. 2 HEATER CYCLE ON ACTUAL

(_)FOR_VARD BATTERY NO. I HEATER CYCLE (UNIT l) ON ------NO DATA (_)FORWARD BATTERY NO, 2 HEATER CYCLE OFF (_)FORWARD BATTERY NO, l HEATER CYCLE (UNIT I) OFF ----_ EXPECTED NOMINAL

(_)FORWARO BATTERY NO, l HEATER CYCLE ACCEPTABLE LIMITS 32 fli!!iiii!iiiiii_!_ii_iiiiiiiiii_i_i_i_!i!i!il!iii!i!i!iiiii!ii_iii_i!iiiiiiiiiii_iiiii!iiiiiilliiiiiiiii!!iiiii_iiiiiiiili_iii!iiiiiiiiiii!i!iiiiili_iiiiiiiiiii!ii!iiiiiiiiiiiii

_ iii_,_i'_iiii_iiiiiiii:,ii;i'Jii_:ii!iiiiii_:!ii!_i::i_!_!;iCii::i;!::Yiiii::!;!;i;_i;!;!;i!_:_!::i_!::!_;:iiC!ii::ii!;!;!i:'_iiiiiiiiiiiii!iiiiiiili!iiiiiiii!iiiiiii!i!iiiiiii_iiiiiiI_i;i_i_iiiiiiiii!_ii!iiiilJiiiiiiiiii_i_iii!ii!ii_i_!i!iiiiii_i_i_!iiiii!iiii_i_i_iiiiiiiiii_iJiiiiil;ii_iiiiiiiiiiiiiiiiiiiii!iNiiii .. _ i_ii_ili:ii:i:iil 24 I 22 i 25

...... (_ . _.-. '-'}"t-E:t-Ed--_--E:}-- U'_-:I- ;-:4-

u 5

o .... -I O I 2 3 4 5 6 7 8 9 10 II 12 13 14 15 16 17 18 19 20 21 22 23 24 _5 26 27 28 29 RANGE TIME, 1000 SECONDS

..... O 00:33:20 01:6:400I I 01:40:00I I 02:] I3:20 .....02:46:40 03:20:00 03:5:2013 I 04:6:402I I 05:00:00i I 05:3:20/ I 06: OI6 :40 ' 06:40:00I i 07: I 3:2O ' 07:6:404 I -00:16:40 DO:]G:40 00:50:00 01:23:20 01:56:40 02:30:00 03:03:20 03:36:40 04:IO:O0 04:43:20 05:I6:40 05:50:00 00:23:20 06:56:40 07:30:00 08:03:20 PAMGE TIME, HOURS:MINUTES:SECONDS

Figure 13-I. S-IVB Stage Forward Battery No. 1 Voltage and Current

C) TRANSFER TO INTERNAL _) S-IVB ESC (_) ENGINE CUTOFF AND RANGE SAFETY NO. 2 OFF (_) PU INVERTER AND DC POWER OFF

_) PU INVERTER AND DC POWER ON ACTUAL

_)PU MIXTURE PATIO 4.5 ON NO DATA (Z) PU pROGRAMMED MIXTURE PATIO OFF (_) S-IVB ECO _ EXPECTED NOMINAL

(_) FORWARD BATTERY NO. 2 DEPLETION ACCEPTABLE LIMITS 32 liiiiiiiiiiiiiiiiiiiiiii...... i:i;ilililili:ili:il . 28

_2G: ...... 24! I I 22 10_

" 6i ® _ ® (Z L -- _ -,_ _..__,

_ 2_ ' ] o: i -1 O 1 2 3 4 5 6 7 8 9 I0 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28 29 RANGE TIE, IO00 SECONDS

.... ' ' ' 2o ' ' ' ' ' ' ' 2 ...... ' ' i _ , _ , 0 00:33:20 01:06:40 01: ._ :OO 02:13:20 02:46:40 03:20:00 03:o3:20 04:26:4D 05:00:00 05:33:20 06:06:40 06:40:00 07:13:20 07:46:40 -00:16:40 00:16:40 00:50:00 01:23:20 01:56:40 02:30:00 03:03:20 03:36:40 04:10:00 04:43:20 05:16:40 05:50:00 06:23:20 06:56:40 07_30:00 08:03:20 RANGE TIME, HOURS:MINUTES:SECONDS

Figure 13-2, S-IVB Stage Forward Battery No. 2 Voltage and Current

13-4 (_ I[RANSft R 1(I INTI RNAt

(_)Afl RATTtRY No. ] IIIAIIN (YEll (10_11 P) 011 ---- ACTUAL

I_ S-IVD ESC

('_S IVD E{:O ------NO DATA

f_AFI BAEILRY IUATED CYC_I EXPECTED NOMINAL

ACCEPTAOLE LIMITS (_ CONIROL VALVLS CYCt IN6 (1YPICAL) _ I_) PASS[VATIUN ENADEE 32

30 _.... _ ;;5:" ' :';'::::;'::: :':::_::5::::;:6::: :;:0:::.::::: " ::::::::::: :::::::::""

:...:...... ::::::2::::: :.:-:.:+:: _:':::'::::::... f:;:::::::::: ...... :.:.:.:.:.:,:.

+ 04_:++_:":i:+:++ _i _ _ i_! +_:..'.'.::!:i:...... _ _ +::+::+::+::ii+i.::+::+i+::+i+i+il.:.:.:.:.:.:+_ _...... _+: :_:+:i:i:+::.+.:;:+:i:+:!:+:+:,:i:!:+:+:!++:!:++:+:!:i_i__ :i:!:i:+:!:+:._++:+:+:!:!:+i:+:i:i:+:+:!_i;_...... '...... +imm:_: _!_ _......

60 f

o n _..J' ..... ------...... -I 0 I 2 3 4 5 6 7 8 9 10 11 12 13 14 I5 16 17 18 19 20 21 22 23 24 25 26 27 28 29

RANGE TIME, 1000 SECONDS

, 0, , 00:3:2031 , 0]: 0I6:40 I 01:4 I0 :O0 I 02:],3;20 ...... 02:46:40 03:B0:O4 03:53:20 04:26:40 05:00:00 05:3:203I ' 06:06:40I I 06:0:004I ' 07:1:20I3 ' 07:46:40I ' -00:16:40 00:16:40 00:50:00 01:23:20 01:56:40 02:30:00 03:03:20 03:36:40 04:10:00 04:43:20 05:16:40 05:50:00 06:23:20 06:56:40 07:30:00 O4:03:20

RANGE TIME, HOURS:MINUTES:SECONDS

Figure 13-3. S-IVB Stage Aft Battery No. l Voltage and Current

C) TRANSFER TO INTERNAL

([_) LOX & LH 2 CHILLDOWN PUMPS OFF (_) AUXILIARY HYDRAULIC PUMP OFF -- ACTUAL (_) AUXILIARY HYDRAULIC PUMP CYCLE NO DATA (3) AUXILIARY HYDRAULIC PUMP ON

--_ EXPECTED NOMINAL (_) fOX & LH 2 CHILLDOWN PUMPS ON

(Z) NO LOAD, LIMITS DO NOT APPLY ACCEPTABLE LIMITS _) HEATER CYCLES 64

62 i_ __) _._. _F_':_._ / I _._,_,_,_ 5D .... i_:" i:::1 _II':! i!: _i:I!_:?,: •( :i :::: I .i; !2:ii!:i ::: Ji:.:_t__L_ 2

_ i : }: ! :lii [iiii 52 . , , l _=_oI_ I,_I I|II I !! I I<_']_I_ I l'J',l] i I _ I | 1| | I | I | °o!--!

RANGE TIME, 1000 SECONDS

I I I I 1 I I I I I I I I I I I I I I I I [ 1 1 1 I I 1 I I I O 00:33:20 01:06:40 01:40:00 02:13:20 02:46:40 03;20:00 03:53:20 04:26:40 05:00:OO 05:33:20 06:06:40 u6:40:00 07:]3:20 07:46:40 -00:]6:40 00:]6:40 00:50:00 01:23:20 O] :56:40 02:30:00 03:03:20 03:36:40 04:10:00 04:43:20 05:16:40 06:60:00 06:23:20 06:56:40 07:30:OO 08:03:20 RANGE TIME, HODRS:MINbTES:SECONDS

Figure 13-4. S-IVB Stage Aft Battery No. 2 Voltage and Current

13-5 The three 5-vdc and seven 20-vdc excitation modules all performed within acceptable limits. The LOX and LH 2 chilldown inverters that furnish power- to the LOX and LH2 recirculation pumps performed satisfactorily and met their load requirements.

Performance of the EBW circuitry for the separation system was satisfactory. Firing units charge and discharge responses were within predicted time and voltage limits. The command destruct EBW firing units were in the required state of readiness if vehicle destruct became necessary.

13.5 INSTRUMENT UNIT ELECTRICAL SYSTEM

All battery voltages and temperatures increased gradually from liftoff as expected. All battery voltages remained within normal limits. Battery currents remained normal during launch and coast periods of flight. Battery power consumption and estimated depletion times are shown in Table 13-4. Battery voltages, currents, and temperatures are shown in Figures 13-5 through 13-7.

The 56-vdc power supply maintained an output voltage of 55.7 to 56.6 vdc, well within the required tolerance of 56 ±2.5 vdc.

The 5-volt measuring power supply performed nominally, maintaining a constant voltage within specified tolerances.

Switch selector, electrical distributors, and network cabling performed nominally.

Table 13-4. IU Battery Power Consumption

RATED POWER CONSUMPTION* ESTIMATED* CAPACITY PERCENT OF LIFETIME BATTERY (AMP-HRS) AMP-HRS CAPACITY (HOURS)

6DIO 350 181.2 51.8 18.9

6D30 350 235.2 67.2 14.4

6D40 350 337.1 96.3 I0.I

*Based on available flight data to 35,214 seconds (09:46:54).

13-6 t:I,LILI11[!iiilI_[!!i!_'I_'I_))_I_ii_ JT

;:lJlll llil[i 35111140 IIIIII) _3olltl, ILlllll

] _I° -

320_ I ' m t_ooFI2D° _31Oliii33011III _3oouIA_ LLL 'II 13o

0 0.5 1 2 4 6 8 lO 12 14 16 18 20 22 24 26 28 30 32 34 36 38 _OPERATING LIMIT RANGE TIME, lO00 SECONDS

--- PREDICTED l i t I i i _ i i .J _ACTUAL 00:33:20 02:46:40 05:00:00 07:1:20 09:26:40 01:40:00 03:53:20 06:06:40 08:20:00 10:33:20 RANGE TIME, HOURS:MINUTES:SECONDS

Figure 13-5. Battery 6DIO Voltage, Current, and Temperature I"L'k l 29 > i, __,_: iiiili!iilii!!ilii!_iiiIiiiil)i!i!iiIi -28_II:;_[I:III . : :_ i_!H:ii_i)itiiilH!iiii!it ": ii!ii!ii:i!Iii_i)i)i!i)iii!H

lili_ 25 I II]i [L Li il

_351 I

_ 4°2530 _. 'i -_'. I! ii

330 1

320 I 31o111i, ! IIIII I i I _120_ 'fill : :_.'._ )F!!°_ i l]lli _o°

0 0.5 2 4 6 I0 12 14 16 18 20 22 24 26 28 30 32 34 36 38 RANGE TIME, lOOO SECONDS OPERATING LIMIT --- PREDICTED ' ' ;6 ' oo ' ; ' ' ' ACTUAL 00:33:20 02: :40 05: :00 07:3:20 09:26:40 01:40:00 03:53:20 06:06:40 08:20:00 I0:33:20 RANGE TI ME, HOURS :MINUTES :SECONDS

Figure 13-6. Battery 6D30 Voltage, Current, and Temperature

13-7 31

29 ..... :i: i::iii ! iii!i!Hi_iiiHiiiiiiliiiiiilHiiii!ii]iiii_i_i !!_iL_i!_iiii1_i_i_ii_ii!_i!iii!_i_i_ii_i_i_iAi!i_i!_ii_A!_!ii_i_!_i_i_i_iiiiii > _::,I I_I ':__I:,': I::T_I' _ItI::I!!ii11:,',i':i!Ii!::!ili',_,ii!li!!':Ii!!iliii!:ilil!':I:_,i_iiI!ililiiiii !i:.H_ '11I,_iI:,_if_::.1 II i_I_',ti::I!I',',11':i::HIiii!Iiiiii!H!iiii:_Ii',i!iiIiiii_,ifililiI',ii!'_iH!iiiiiiliii',i > 27 i:ii iii:iiii! ; if: i:i:i!:.iii.; :: i:

_. 35 -._ ._J,,,,m ,,("_'_ 3O

15 OFF SCALE TO 340°K (152.6°F) "_ 33o I l! -130 -120 u_ 320 J o -llO _31o I I -100 I-- -90 _- _; 300 J .80 -70 _: 290 -- -60 280 '#I "50 0 0.5 1 2 4 6 8 I0 12 14 16 18 20 22 24 26 28 30 32 34 36 38 OPERATING LIMIT RANGE TIME, lOOO SECONDS ----- PREDICTED ______ACTUAL 00:33:20 02:46:40 05:00:00 07:13:20 09:26:40 Ol:4O:OO 03:53:20 06:06:40 08:20:00 I0:33:20 RANGE TIME, HOURS:MINUTES:SECONDS

Figure 13-7. Battery 6D40 Voltage, Current, and Temperature

13-8 SECTION14

RANGESAFETYAND COMMANDSYSTEMS

14.1 SUMMARY

Data indicated that the redundant Secure Range Safety CommandSystems (SRSCS) on the S-IC, S-II and S-IVB stages were ready to perform their functions properly on commandif flight conditions during the launch phase had required vehicle destruct. The system properly safed the S-IVB SRSCSon a commandtransmitted from Bermuda (BDA). The performance of the Commandand Communications System (CCS) in the Instrument Unit (IU) was satisfactory, except for the Radio Frequency (RF) problem noted in paragraph 19.4.3.2. 14.2 SECURERANGESAFETYCOMMANDSYSTEMS

Telemetered data indicated that the commandantennas, receivers/ decoders, Exploding Bridge Wire (EBW) networks, and destruct controllers on each powered stage functioned properly during flight and were in the required state of readiness if flight conditions during the launch phase had required vehicle destruct. Since no arm/cutoff or destruct commandswere required, all data except receiver signal strength remained unchanged during the flight. Power to the system was cut off at 723.5 seconds by ground commandfrom BDA, thereby deactivating (safing) the system. Both S-IVB stage systems, the only systems in operation at this time, responded properly to the safing command.

Radio Frequency (RF) performance aspects of the system are discussed in paragraph 19.4.3.1. 14.3 COMMANDAND COMMUNICATIONSYSTEM

The commandsection of the CCSoperated satisfactorily except for the RF problem noted in paragraph 19.4.3.2. Twenty commandswere initiated by Mission Control Center - Houston (MCC-H) for transmission via the Gold- stone (GDS) Wing Station, as shown in Table 14-I. The last II commands were initiated with the ground station in the Message Acceptance Pulse (MAP) override mode. The MAPoverride mode was necessary because the telemetry data was noisy and the Address Verification Pulses (AVP's) and Computer Reset Pulses (CRP's) could not be detected at the ground

14-I station. A total of 50 commandwords were attempted by the GDSWing Stati on. The commandat 19,033.7 seconds (05:17:13.7) to switch the CCScoaxial switch to the low-gain directional antenna position was unsuccessfully transmitted four times. The commandwas not received by the onboard equipment because the uplink subcarrier was not in lock. Upon completion of the automatic commandcycle (the ground station is set up to automatically transmit the commandword four times or until the AVP's and CRP's are received), a terminate commandwas issued to reset the commandsystem and the switch commandwas again attempted at 19,062.3 seconds (05:17:42.3). During this second transmission, the ground computer failed to capture the AVP's and CRP's, resulting in the command being repeated three times. The verification pulses were missed because

Table 14-I. Command and Communication System GDS Commands History

RANGE TIME NUMBER OF SECONDS 'HRS:MIN:SEC COMMAND WORDS REMARKS

17,466.6 04:51:06.6 T8 Initiated Accepted 17,770.9 04:56:10.9 Begin Environmental Control System (ECS) Experiment 1 Accepted

18,502.2 05:08:22.2 Engine He Control Valve Enable 6 Accepted 19,033.7 05:17:13.7 Set Antenna Low Gain 4* Uplink Subcarrier Out-of-Lock 19,051.9 05:17:31.9 Terminate 1 Accepted

19,062.3 05:17:42.3 Set Antenna Low Gain 4* Accepted 27,367.7 07:36:07.7 Set Antenna High Gain 2* Accepted

32,019.3 08:53:39.3 CCS Transponder Disable 4 Noisy Telemetry** 32,066.6 08:54:26.6 CCS Transponder Disable 4 Noisy Telemetry** 32,601.4 09:03:21.4 CCS Transponder Disable 3 Accepted (MAP Override) 32,669.5 09:04:29.5 CCS Transponder Enable 3 Accepted (MAP Override) 33,825.1 09:23:45.1 Set Antenna Omni 1 Accepted (MAP Override) 34,000.1 09:26:40.1 Set Antenna Low Gain 1 Accepted (MAP Override) 34,105.3 09:28:25.3 CCS Transponder Disable 3 Accepted (MAP Override)

34,160.0 09:29:20.0 CCS Transponder Enable 3 Not Transmitted by Ground Station

34,234.5 09:30:34.5 Set Antenna Omni Acceptance Status Unknown (MAP Override)

34,312.0 09:31:52.0 Set Antenna High Gain Acceptance Status Unknown (MAP Override) Antenna Omni 34,419.9 09:33:39.9 Set Acceptance Status Unknown (MAP Override)

34,530.0 09:35:30.0 CCS Transponder Disable Acceptance Status Unknown (MAP Override)

34,554.9 09:35:54.9 CCS Transponder Enable Accepted (MAP Override)

*One word is normally required to switch antennas. These commands were repeated due to the uplink being out of lock or missed verification pulses at the ground station because of noisy telemetry.

**Only Mode Words Transmitted

14-2 of noisy telemetry due to low downlink signal strength. Acceptance of the commandwas verified by an increase in signal strength and by the antenna position measurement (K132-603) indicating the CCScoaxial switch was in the low-gain antenna position.

Noisy telemetry resulted in a repeated commandat 27,367.7 seconds (07:36:07.7) to transfer the CCS coaxial switch to the high-gain antenna position. The commandwas repeated once before the ground computer detected the acceptance pulses and terminated transmission of the command.

Transmission of the CCSdisable commandwas unsuccessful when attempted at both 32,019.3 seconds (08:53:39.3) and 32,066.6 seconds (08:54:26.6) due to noisy telemetry. The noise prevented the ground station from detecting the AVP's and CRP's. Therefore, acceptance of the mode word could not be verified. The mode word was transmitted eight times before the MAPoverride mode was selected and the complete commandtransmitted (one mode and two datawords). The commandwas accepted on this thi.rd attempt at 32,601.4 seconds (09:03:21.4).

The commandto enable the CCSat 34,160.0 seconds (09:29:20.0) was not transmitted by the ground station because the 70-kilohertz subcarrier was off. This meant that the CCSdownlink was inhibited from 34,105.3 seconds (09:28:25.3) (CCS disable command)until the enable commandtransmitted at 34,554.9 seconds (09:35:54.9) was accepted. This mode was verified by the signal strength level during this period (see paragraph 19.4.3.2). Since the downlink was inhibited during this period, no AVP's and CRP's were received for the antenna switching commandsand the disable commandwas not transmitted during this period.

The acceptance status of the commandstransmitted during the period in which the CCSwas inhibited is unknown except for the two commandsto switch to the omni antennas. One or both of these commands, 34,234.5 seconds(09:30:34.5) and 34,419.9 seconds (09:33:39.9), were accepted by the onboard system because measurement K131-603 indicated the system was on the omni antennas when the downlink signal returned at 34,555 seconds (09:35:55).

14-3/14-4

SECTION15

EMERGENCYDETECTIONSYSTEM

15.1 SUMMARY

The performance of the AS-506 Emergency Detection System (EDS) was normal, and no abort limits were exceeded.

15.2 SYSTEMEVALUATION

15.2.1 General Performance

The AS-506 EDSconfiguration was the same as on AS-505. All launch vehicle EDSparameters remained well within acceptable limits during the AS-506 mission. EDSrelated sequential events and discrete indications occurred as expected.

15.2.2 Propulsion System Sensors

The performance of all thrust OK sensors, which monitor engine status, was nominal insofar as EDSoperation was concerned. The associated voting logic was also nominal. S-II and S-IVB tank ullage pressures remained within the abort limits, and displays to the crew were normal.

15.2.3 Flight Dynamics and Control Sensors As noted in Section II, none of the triple redundant rate gyros gave any indication of angular overrate in the pitch, yaw, or roll axes. The maxi- mumangular rates were well below the abort limits. The roll rate abort limit is 20 deg/s; a switch selector commanddeactivated the overrate automatic abort and changed the pitch and yaw rate abort settings from 4 deg/s to 9.2 deg/s at 134.8 seconds. The maximumangle-of-attack dynamic pressure sensed by a redundant Q-ball mounted atop the escape tower was 0.28 N/cm2 _0.4 psid) between 89 and 91 seconds. This pressure was only 12.5 percent of the EDSabort limit of 2.2 N/cm2 (3.2 psid).

15-I/15-2

SECTION16

VEHICLEPRESSUREENVIRONMENT

16.1 SUMMARY

The S-IC stage base pressure environments were monitored by two heat shield differential pressure measurements. S-II stage base pressure environments were monitored by two absolute pressure measurements on the heat shield and one on the thrust cone. The flight data were generally in good agreement with the postflight predictions and compared well with previous flight data. The pressure environments were well below design levels.

There was no instrumentation provided on the AS-506 vehicle which would permit a direct evaluation of the surface and compartment pressure environments. One internal ambient pressure measurement located on the S-II forward skirt was used to calculate the pressure loading acting on that area and agreed with predictions and previous flight data. 16.2 BASEPRESSURES

16.2.1 S-IC Base Pressures

The S-IC stage base heat shield pressure loading was recorded by two differential pressure measurements. Both measurements show good agreement with previous flight data as shown in Figure 16-I. Pressure loading is the difference between internal and external pressures (Pint-Pbase) defined such that positive loading is in the burst direction. The heat shield loadings were well within the 1.4 N/cm2 (2.0 psid) design pressure loading. 16.2.2 S-II Base Pressures

The S-II stage base heat shieid and thrust cone pressure environment was recorded by two absolute pressure measurements on the heat shield and one absolute pressure measurement on the thrust cone.

Except for the absence of a more significant drop in measured aft face pressure at S-II Center Engine Cutoff (CECO), Figure 16-2 shows good agreement between the postflight predicted and AS-506 flight heat shield aft face static pressure history. It is seen that the AS-506 pressure falls within the AS-501 through AS-505 data band. The predicted pressure drop after S-II CECOis based on the computed total pressure loss resulting

16-I ALTITUDE, n mi 5 I0 15 20 25 30 0.4 l l

E U - 0.4 "_ (__ Z zz 0.2 ---- AS-505 ._r- AS-506 O.2-J __J u_ n_ STA 2.81 m L_._ r_r_ I,I.Q c_c__ ::) I m ..... ___._ _II,(J.63 in.) V') 4.J ct3 4-} :__. -- 0 Cz') (-" LI.J .r-- r,,.," r_ I.I-I °r" C_C_. CL -.,-._ - -0.2 -0.2

0.4

O,J E - 0.4 U C..') _ zz 0.2

E) _J /" _V_------/--AS-506AS-504 D0047-106 - 0.2--_ ._.J _,_ Ill STA 2.81 m l.a.l ..J_ i.I.ir'l r,,." r_ rY" i__ (110.63 in.) I _° _ " I L°''-o_-_ I (I')4.-) _'_ 0 U (.I) ,c" (._ e-" LI.J • I" r_r_

- -0.2

-0.2 0 10 20 30 40 50 60 ALTITUDE, km

Figure 16-I. S-IC Base Heat Shield Pressure Loading

from the reverse flow passing through a shock wave above the nozzle lip of the inoperative center engine. Based on AS-505 flight data, a somewhat smaller but still measurable drop was expected for the D158-206 measurement. The further pressure reduction occurring after Engine Mixture Ratio (EMR) shift is predicted from the reduction of the maximum pressure in the J-2 engine exhaust plume interaction regions.

Figure 16-3 shows the static pressure variation with range time on the forward face of the base heat shield. It is seen that the AS-506 measured static pressure on the forward face of the heat shield, while within design limits, exceeds the postflight prediction and was

16-2 S-II ESC X_ S-II CECO X_7 S-II OECO _7 S-II ESC V S-II SECOND PLANE X_ EMR SHIFT SEPARATION 7 S-ll SECONDPLANE SEPARATION TRANSDUCER(D158-206) VEH STA 40.9 m (1608 in.) "_/_-_" PREVIOUS X_7 S-II CECO RADIUS 1.91 m (75 in.) /_-'_L._/_\ I AZIMUTH35169 __ FLIGHT DATA _7 EMR SHIFT I FLIGHT DATA PREVIOUS __\_"_%v / - - -- - POSTFLIGHT PREDICTION _7 S-II OECO 0.07 I I I - 0.I0 FLIGHT DATA _ = 9o_''--L_ _ : o° ---- POSTFLIGHT PREDICTION - /" 0.06 0.07 - 0.I0 I_//TRANSDUCER (D150-206) --0.08

0.06 0.05 I-- \ _r/ -VEH STA 40.9 m (1610 in.) I .)<_ _Y ->{ RADIUS 1.85 m (73 in.) -- 0.08 ! o=9_ _'- _= _ AZIMUTH 350 deg __ o.o4 i I i -0.O6 & 0.05 i i I I -0.06 . _, o.o3 I I I c,3 0.04 #. I r E -0,04 _ D.. u_ I r I 0.03 o.o2 -0.04 _- ! I l -- 0.02 =< I 0.02 0.01 I

- 0.02 O II 1"_- 0.01 -0 , V 'V o 150 200 250 300 350 400 450 500 550 600 150 200 250 300 350 400 450 500 550 600 RANGE TIME, SECONDS RANGE TIME, SECONDS

Figure 16-2. S-II Heat Shield Aft Face Pressure Figure 16-3, S-II Heat Shield Forward Face Pressure approximately 30 percent higher than that measured during the AS-501 and AS-502 flights. No pressure measurement was available at this exact location during the AS-503 through AS-505 flights. This condition is believed to be a localized effect due to variable leakage through the J-2 engine nozzle flexible curtains.

Figure 16-4 shows the AS-506 static pressure variation on the thrust cone, The measured AS-506 thrust cone static pressures agreed well with predicted values and with previous flight data.

_7 S-II ESC

V S-II SECOND PLANE SEPARATION

_7 S-II CECO I FLIGHTPREVIOUSDATA _7 EMR SHIFT FLIGHT DATA POSTFLIGHT PREDICTION _7 S-II OECO 0.07 I0

0.06 VEH STA 43.1 m (1698 in.) AZIMUTH 2 deg O8 TRANSDUCERRADIUS 3.30 (D187-206)m (130 in.)_ ___ m 0.05

E u 06 "_ z 0.04

l,J r_ r_ 0.03 ULJ 0.04 r-/ r_

0.02 rY_ o.o 0.02 ! I o !

-o.ol 7 R7 150 200 250 300 350 400 450 500 55O 600

RANGE TIME, SECONDS Figure 16-4. S-II Thrust Cone Pressure

16-4 16.3 SURFACE PRESSURE AND COMPARTMENT VENTING

16.3.1 S-IC Stage

There was no instrumentation on the S-IC stage for evaluation of the surface and compartment pressure environments.

16.3.2 S-II Stage

Other than the internal ambient pressure measurement (D163-219) located on the forward skirt, there was no instrumentation on the S-II stage for evaluation of the surface and compartment pressure environments. A calculated pressure loading (Pint-Pext) on the forward skirt was obtained by taking the difference between the predicted external pressure values and the internal pressure (assumed uniform), which was measured at vehicle longitudinal station 62.2 m (2449 in.) and peripheral angle of 191 degrees (see Figure 16-5). The AS-506 flight data (calculated) show the same trends and are in good agreement with the postflight predictions and previous flight data.

16-5 LIFTOFF _7 MAXIMUM DYNAMIC PRESSURE

MACH 1 _7 S-IC/S-II Ist PLANE SEPARATION

INTERNAL MEASUREMENT: D163-219

POSTFLIGHT PREDICTION VEHICLE STATION: 62.2 m (2449 in.) AS-506 FLIGHT DATA (CALCULATED) AZIMUTH ANGLE: I DATA FROM PREVIOUS FLIGHTS _= 191 degrees

DESIGN LIMIT (STATIC PRESSURE) VARIES FROM4.34 N/cm 2 (6.3 psid) TO 4.69 N/cm 2 (6.8 psid) DEPENDIN_ ON VEHICLE STATION

/- --7

I _=0 = 90 !

E o 4 i --6 I i --5 _-_ X 4) GJ X r_

' 3 ! ! c-- ore- _4 t- i .r-- v I v i z I 'I] Z _3 c: 2 r-% 0 .=J m 0 i,i ,,, _2 _" A il _0 O0 1,1 I| I fs I r_ rv" t --I : i _,

I 0 t,

' - ] ,,2,, \i/ r I 0 25 50 75 1O0 125 150 175

RANGETIME, SECONDS Figure 16-5. S-II Forward Skirt Pressure Loading

16-6 SECTION 17

VEHICLE THERMAL ENVIRONMENT

17.1 SUMMARY

The AS-506 S-IC base region thermal environments have similar magnitudes and trends as those measured during previous flights. Maximum values of total heating and gas temperature were recorded at approximately 20 kilometers (10.8 n mi) altitude with maximum values of 25 watt/cm2 (22.2 Btu/ft2-s) and 1200°K (1695°F), respectively.

In general, base thermal environments on the S-II stage were similar to those measured on previous flights and were well below design limits. However, the heat shield aft radiation heating rates were approximately 20 percent higher than the maximum values measured during previous flights.

Flow separation was observed (ALOTS film) to occur at approximately 116 seconds range time. Aerodynamic heating environments were not measured on AS-506.

17.2 S-IC BASE HEATING

Thermal environments in the base region of the S-IC stage were recorded by two total calorimeters and two gas temperature probes which were on the heat shield at the locations shown in Figure 17-I. Data from these instruments are compared with the AS-502 through AS-505 flight data band (Figures 17-2 and 17-3) and are shown versus altitude to minimize trajectory differences. AS-501 flight data, which showed less severity than subsequent flight data because of flow deflector effects, are not shown.

As shown in Figures 17-2 and 17-3, the AS-506 S-IC base heat shield thermal environments have similar magnitudes and trends as those measured during the previous flights. Maximum values of total heating and gas temperature data were recorded at approximately_ 20 kilometers (10.8 n mi) with maximum values of 25 watt/cm2 (22.2 Btu/ftZ-s) and 1200°K (1695°F), respectively. Center Engine Cutoff (CECO) on AS-506 produced a spike in the data with a magnitude and duration similar to previous flight data at CECO. The AS-506 gas temperature data are similar to previous flight data. However the AS-506 and AS-505 gas temperature data do not show the decrease between 4 and 9 kilometers (2.2 and 4.9 n mi) which the AS-502 through AS-504 flight data indicated.

17-I Ambient gas temperatures inside the engine cocoons remained within the band of previous flight data.

POS IV _ POS III

C149-I06 TOTAL CALORIMETER C050-I06 GAS TEMPERATURE (0.64 cm (0.25 f FT OF SURFACE) in.) POS I POS II C026-I06 TOTAL CALORIMETER \ FIN A C052-I06 GAS TEMPERATURE VIEW LOOKING AFT (6,35 cm (2.50 in.) AFT OF SURFACE)

Figure 17-I. S-IC Base Heat Shield Measurement Locations

17-2 _o '3_n1V_3dW31 SV9 Jo '3_nlV_3dW31 SV9

(3J c" (./3

_J mr- QJ e (1; S-- e:_ "1--) S----

I E

• (/) (Y3 r_ o o r" CM O CO _o ,_P (3J _Io '3_l_rd3dH31 SV9 E_ °1" _Lr) LL

[_i_!],

s-2%$/n_Q s-_In_Q '31_rd 9NI±V3H lV±Ol '31_rd 9NIIV3H 7VI01

_3 _N mcr_

r..=.. o4.._

'3lW 9NI±V3H ]VIOl ' 3.l.V'd 9NI/V3H ]VlO.l.

C_ •f,-- IJ._

17-3 17.3 S-II BASE REGION ENVIRONMENT

The S-II base heat shield flight environments were, in general, in good agreement with previous flight data and were well below design limits. S-II heat shield aft face convective heating rates, aft radiation heating rates, and base gas temperatures are presented in Figures 17-4 through 17-6, along with previous flight data and postflight predictions. The postflight predicted heat shield convective heating rates are based on hot flow 1/25 scale S-II stage model test data. Postflight predictions for the other two figures were accomplished by the same analytical methods that have been described in previous flight evaluation reports.

sll Esc V sIf cEco V S-II OECO

S-II SECONDPLANE SEPARATION _ EMR SHIFT

PREVIOUS FLIGHT TRANSDUCER (C722-206) __ I DATA VEH STA 41.4 m (1630 in.) _ _J_._ RADIUS1.9 m(75 i,.]_____---_. FLIGHT DATA AZIMUTH 83 deg --_-_- _ =.eX_ PREDICTION

! ! E u l & 4- "'_V 'Y ] 2 rr_ l w F-- F--

z 1 Z

i,i I I,i -r -r

-I -I

! ! VV V 150 2O0 250 300 350 400 450 500 550 6O0

RANGE TIME, SECONDS

Figure 17-4. S-II Heat Shield Aft Face Heat Rate

17-4 V S-II ESC V S-II SECOND PLANE SEPARATION V S-II CECO I PREVIOUSFLIGHT DATA V EMR SHIFT FLIGHT DATA PREDICTION S-II OECO , I I I I I

_j_ TRA_sDUCER(C692-206) - ,_ --VEH STA 41.4 m (1630 jn X RADIUS 2.54 m (I00 in ) o 9o° I _ = oO--AZIMUTH 0 deg T c_ I I E (J

+_ 2 4_

w --_-Z ...... _ _

Z Z b-"

w / -r 0 0

-1 -I

250 300 350 400 450 500 550 600 RANGE TIME, SECONDS

Figure 17-5. Heat Shield Aft Radiation Heat Rate

Figure 17-5 shows that the incident radiative heat flux during the AS-506 flight was greater than predicted and approximately 20 percent higher than the maximum values measured during flights AS-501 through AS-505. The most probable cause for this increase is engine misalignment or engine gimbaling, neither of which are accounted for in the postflight prediction of the incident radiative heat flux.

There were no measurements of structural temperatures made on the AS-506 S-II stage base heat shield. To evaluate the structural temperatures experienced on the aft surface of the heat shield, a maximum postflight

!7-5 S-II ESC S-II OECO V S-II CECO X_

V S-II SECOND PLANE SEPARATION EMR SHIFT

NOTE: THE PROBES EXTEND I FLIGHTPREVIOUSDATA APPROX. 0.05 m (2 in.) ABOVE

.... GAS RECOVERY THE HEAT SHIELD AFT SURFACE_ TEMPERATURE TRANSDUCER (C731-206) , /,_"_"_ FLIGHT DATA VEH STA 41.4 m (1630 in.) PROBE TEMPERATURE RADIUS 1.91 m (75 in.) __

_----PREDICTION AZIMUTH 99 deg_ _._ PROBE TEMPERATURE _:'9_ 1200 1600 1100

I000

1200 900

40O

300 0 200

I00

0 150 200 250 300 350 400 450 500 550 600

RANGE TIME, SECONDS Figure 17-6. S-II Base Gas Temperature predicted temperature was calculated for the aft surface using base heating rates predicted for the AS-506 flight. The predicted maximum postflight temperature was 818°K (IOI4°F) which compared favorably with maximum postflight temperatures predicted from previous flights, and was well below the maximum design temperatures of I066°K (1460°F) for the no-engine-out case and III6°K (1550°F) for the one control engine-out case.

17-6 The effectiveness of the heat shield and flexible curtains as a thermal protection system was again demonstrated on this flight as on previous flights by the relatively low temperatures recorded on the thrust cone forward surface. The AS-506 maximummeasured thrust cone forward surface temperature was 266°K (20°F), essentially equal to that recorded during previous flights. The measured temperatures were well below design values and in good agreement with postflight predictions. 17.4 VEHICLEAEROHEATINGTHERMALENVIRONMENT

17.4.1 S-IC Stage Aeroheating Environment

The aerodynamic heating environments were not measured on the AS-506 S-IC stage. However, flow separation was measurable from flight optical data (ALOTS film) and was observed to occur at approximately the same time as on AS-505, 116 seconds as shown in Figure 17-7. The effects, of CECOon the separated flow region during AS-506 flight were the same as observed on previous flights. It should be noted that at higher altitudes, the measured location of the forward point of flow separation is question- able because of loss of resolution in the flight optical data.

6O NOTE: FORWARDLOCATION OF FLOW SEPAP_ATION MEASURED FROM AS-506 ALOTS 70 MM FILM 5O 2000 V S-IC CECO

E 40 1500

0 I00 1 I0 120 130 140 150 160 RANGE TIME, SECONDS

Figure 17-7. Forward Location of Separated Flow

17-7 17.4.2 S-II Stage Aeroheating Environment

There were no aerodynamic heating environments measured on the S-II stage; however, postflight predicted temperatures were determined based on the actual AS-506 trajectory and thermal models used in previous flight evaluations. All postflight predicted temperatures were well below the design limits and within the band of previous flight data.

17-8 SECTION18

ENVIRONMENTALCONTROLSYSTEM

18.1 SUMMARY

The S-IC canister conditioning system and the aft environmental conditioning system performed satisfactorily during the AS-506 countdown.

The S-II thermal control and compartment conditioning system apparently performed satisfactorily since the ambients external to the containers were nominal and there were no problems with the equipment in the con- tai ners.

The Instrument Unit (IU) Environmental Control System (ECS) exhibited overall satisfactory performance for the duration of the IU mission. Coolant temperatures, pressures, and flowrates were continuously maintained within required ranges and design limits. One deviation from specification was observed. The inertial platforf_ gas bearing differential^pressure drifted above the I0./ N/cmz (15.5 psid} maximum to 11.2 N/cmz (16.3 psid). This drifting phenomenon also occurred on on AS-501, AS-503, and AS-504 and caused no detrimental effect on the mission. 18.2 S-IC ENVIRONMENTALCONTROL

The ambient temperatures of the 4 canisters in the S-IC forward skirt compartment must be maintained at 300 +_II°K (80 +20°F) during equipment operation prior to J-2 engine chilldown and 325 to 278°K (125 to 40°F) during J-2 engine chilldown. No canister conditioning is required after S-IC forward umbilical disconnect.

The canister conditioning system was supplied with air/GN 2 (gaseous nitrogen) at a flowrate of 17.24 kg/min (38 Ibm/min) and a temperature of 299°K (79°F) through the S-IC forward lower umbilical and at a flowrate of 15.42 kg/min (34 Ibm/min) and a temperature of 301°K (81°F) through the S-IC forward upper umbilical during AS-506 countdown prior to J-2 engine chilldown. During J-2 engine chilldown, the flowrate and temperature of the GN2 supplied to the forward upper umbilical was increased to 18.82 kg/min (41.5 Ibm/min) and 311°K (IO0°F), and the temperature of the GN2 supplied to the forward lower umbilical was increased to 314°K (I05°F). No instrumentation was installed in the canisters on AS-506; therefore,

18-I no evaluation of the actual temperatures within the canisters was possible. No failure of any electrical/electronic equipment installed in the canisters was reported.

During J-2 engine chilldown, the thermal environment is at the most critical point. Within this period the ambient temperature in the forward skirt compartment dropped as shown in Figure 18-I. The lowest ambient temperature measured during AS-506 J-2 engine chilldown was 229°K (-48°F) at instrument location C206-120. During AS-506 flight, the lowest temperature recorded was 183°K (-130°F) at instrument location C206-120.

The design requirement for the aft compartment is that the ambient temperature for prelaunch be maintained at 300.0 +8.3°K (80 +I5°F). Aft compartment prelaunch ambient temperatures are shown in Figure 18-2. The lowest prelaunch temperature recorded was 287°K (58°F) at instrument CI07-I15. This low temperature occurred prior to LOX loading and did not cause any problems. Aft compartment ambient temperatures for flight are also shown in Figure 18-2. The lowest temperature recorded was 285°K (54°F) at instrument C203-I15. 18.3 S-II ENVIRONMENTALCONTROL

The engine compartment conditioning system maintained the ambient and thrust cone surface temperature within design ranges throughout the launch countdown. The system also maintained an inert atmosphere within the compartment. There were no thermal control container temperature measurements; however, since the ambients external to the containers were satisfactory and there were no problems with the equipment in the containers, it is assumed that the thermal control systems performed adequately. The ambient temperature near the forward system was 44.5 to 85°K (80 to 152°F) warmer than the extremes of past vehicles due to the increased effectiveness of the foam insulation used on the S-II-6 hydrogen tank forward bulkhead. Foaminsulation reduced heat leakage from the engine compartment and resulted in more uniform temperatures from container to container. 18.4 IU ENVIRONMENTALCONTROL

The IU ECS is composed of a Thermal Conditioning System (TCS) and a Gas Bearing Supply System (GBS). A preflight purge subsystem provided compartment conditioning prior to launch and maintained the compartment temperature within the required 290 to 296°K (63 to 73°F) range.

18.4.1 Thermal Conditioning System

Initial sublimator start-up and sublimator performance parameters during ascent are depicted in Figure 18-3. Immediately after liftoff, the Modu- lating Flow Control Valve (MFCV) began driving toward the full heatsink

18-2 AS-506 FLIGHT DATA 260 I i

,0 -----_ f c208-120, 250

-2O

v 240, h o 0

I--- -40 p- _ 230 \ _- C206-I 20 o_

I--- -6O 220

210 -8O _r.'MINIMUM PERFORMANCELIMIT _ _ _ _r_ .I J

2( -I 80 -160 -I O0 -80 -60 -40 -20

280 AS-506 FLIGHT DATA "40 I I I I

260 \ PREDICTED , f MAXi MUM -0 \

240 _ '. \ "-40 o /

220 _ "- l_ _ ---2 fi --80 200 , _'PREDICTED MINIMUM ."--____'"'-....., __._.__"-_,_J c206-12o" --_..._-:_ _" "-120 180

"-160 160

0 20 40 60 80 I00 120 140 160 RANGETIME, SECONDS Figure 18-I. S-IC Forward Compartment Ambient Temperature

18-3 -_ ' 3_NI_Fd3dH3.L -lo °3_n.1_"d3dH31 c_ ._0 '3_nlD"d3d_31 o c_ o i ._o I i E i _ i It / i ._. I i I I -N-_ !, Ill x I ) ,_. I Illl T I i Itll E I Jill

I/ll w_ I/ll o I _ Illl ! i oO 'T I 'T I j I

li ! I I I I rl I II } s N _//"

"- Ii I oO ( z ", ,/ i/ co ! IIII _ I ( I IIii _ o o _ N N o _ _ o_ "_o ' 3_AlV_3d_31 i, _o ' 3Unl_3d_P_t _ '3_IW3dH31 292 I I I i - 65 ' 291 -M-W CONTROL TEMP C15-601_

290 o . 289 I THERMuAL " 60 2 288 SWITCH _7 WATER VJLVE OPEN •_ CLOSED TEMP 2 I-- _287 _7 FIRST THERMAL SWITCH SAMPLING _: 286 _-WATER VALVE CLOSED- _- -55 _

285 I--

284

283 - 5O u2.5 32.0 SUBLII_ATOR W'ATER 3_ i INLET PRESSURE D43-601 I" \ - 1.5 22 _I.0 _o.5 •= o o 2

lO SUBLIMATOR'HEAT .ESTIMATED 8 --REJECTION RATE - I . 6 I 20,000 / 4 l/ "I0,000"30'000 1 _- 2

1 0 lO0 200 300 400 500 600 700 800 900 lO00 llO0 1200 1300 1400 1500 RANGE TIME, SECONDS _7 , R7 , _ , , , 0:05:00 O:lO:O0 0:15:00 0:20:00 0:25:00 RANGE TIME, HOURS:MINUTES:SECONDS

Figure 18-3. Sublimator Performance During Ascent

position which was reached at approximately 30 seconds. The water valve opened at 181 seconds allowing water to flow to the sublimator. Immed- iate cooling was evidenced by the rapid decline in the coolant fluid temperature. At the first thermal switch sampling, the coolant temperature was still above the actuation point and the water valve remained open. The second thermal switch sampling at approximately 780 seconds resulted in closing of the water supply valve.

Coolant flowrates and pressures were well within required ranges as indicated in Table 18-I.

An after mission experiment was performed in which the water supply valve logic was inhibited (valve closed) to determine the effect of loss of sublimator cooling. This was initiated approximately 5 hours after liftoff. The Methanol/Water (M/W) supply temperature exceeded the maximum scale range of 293°K (68°F) at about 23,200 seconds (Figure 18-4).

The TCS GN2 sphere pressure decay which is indicative of GN2 usage rate was as expected for the nominal case as shown in Figure 18-5.

18-5 Table 18-I. TCS Coolant Flowrates and Pressures

PARAMETER REQUIREMENT MINIMUM MAXIMUM OBSERVED OBSERVED

IU Coolant Flow- 2.18 2.20 2.27 rate F9-602 (9.6) (9.71 (I0.0) m3/hr (gpm)

S-IVB Coolant 1.77 +_.09 1.77 1.82 Flowrate F10-601 (7.8 +_.4) (7.8) (8.2) m3/hr (gpm)

Pump Inlet Pres- 10.83 to LII.72 II .03 11.38 sure D24-601 (15.7 to _17.0) (16.0) (16.51 N/cm 2 (psia)

Pump Outlet Pres- 28.89 to ,33.23 31.03 31.72 sure D17-601 (41.9 to 48.2) (45.O) (46.0) N/cm 2 (psia)

294 M)W C'ONTROL'TEMPEP_TUI_E ('C15'-601') 70 293 I

292

291 65

290

:_ 289 o " /1 I , 6o_ " 288

I-- _ 287 _ 286

285 I-.-

284 L 283

282

281 28O 45

:_0.04 _oo2 I Sl]BLiMAT'ORWATER F'LOWRATE (F'l-6()l)- 0.2 _E 0 ' o N _J % 3.0 I I I I I I I I I I I I 1'-4 ,'_ oo2.5Oo -SUBLIMATORiiiiWATER I LE P SS RE (D4 iil-6 I) o"_ 0 2 4 6 8 I0 12 1416 ]8 20 22 24 26 28 30 32 34 36 38 40 42 44 46 48 50 RANGE TIME, I000 SECONDS

I I t I i I I l I I I I I I ; 0 2:00:00 4:00:00 6:00:00 8:00:00 I0:00:00 12:00:00 14:00:00 RANGE TIME, HOURS:MINUTES _SECONDS

Figure 18-4. TCS Coolant Control Parameters

18-6 23 22 -3200

21 . -3000 20 19 2800 18 2600 17 % -2400 16 NOMINAL EXPECTED DECAY 15 / 2200 14 2000 13 Z -1800 % 12 II 1600

L_J I0 -1400 9 8 -1200

7 I000 6 800 5 4 6OO

4O0

2OO

0 0 2 4 6 8 I0 12 14 16 18 20 22 24 26 28 30 32 34 36 38 40 42 44 46 48 50

RANGE TIME, I000 SECONDS

I I I I I I I I I I I I I I I I:00:00 3:00:00 5:00:00 7:00:00 9:00:00 II:00:00 13:00:00 2:00:00 4:00:00 6:00:00 8:00:00 I0:00:00 12:00:00

RANGE TIME, HOURS:MINUTES:SECONDS

Figure 18-5. TCS GN2 Sphere Pressure (D25-601)

All component temperatures remained within their expected ranges for the duration of the mission (Figure 18-6). The ST-124M internal gimbal (inertial) temperature (C34-603) went below operational temperature range 313°K (I04°F) (marginal operation) at about 4 hours. Lower temperature operation was also observed on AS-504 and AS-505, and is due to a change in internal platform configuration (including axial blower) effect- tive on AS-504 and subsequent. No degradation of platform performance has been noted. The component temperatures all climbed as expected during the ECS experiment and C34-603 went above its upper operating limit 319°K (II5°F) at about 9 hours (Figure 18-6).

18.4.2 Gas Bearing Supply System

The GN2 pressure differential across the ST-124M platform gas bearings drifted from an initial value of 10.48 N/cm 2 (15.2 psid) at liftoff to 11.24 N/cm2 (16.3 psid) at 23,200 seconds (see Figure 18-7). The upper limit of the specification value 10.7 N/cm2 (15.5 psid) was exceeded at about 2500 seconds. The phenomenon of upward drifting of the pressure differential has occurred on AS-501, AS-503, and AS-504 flights. Exten- sive analysis and laboratory testing has indicated that the pressure

18-7 :lo '3_Nl_d3_31 _-Io '3_NL_d3d_31 • . . iiiiii...,...,_,,..,_;.;._;,;iiiilliiii • . o _ i_i:!:_:i:i::i:i:i:i:_ o '_:i::_!:i:i:i:_::i:i:i:!: % • ,,%°• t, 0.o. ":_:':':':':':':":':':'!' "=:"_-';°_ "°°° 1 • • o

iiiiiiiiii_i_i!!i _ *. ee :i:i:i:i:!:_:_:i::::::;!;!_o ".. 5 :i:i:i:i:i:i:i:i::,:.:.:-:o

_vooa c.._ >- ¢'_,,I c'_J :::::::::::::::::::::::::: 0 •".'.'.'-'-','-',; ";':': 0

I .:.:-:-:-:.:-:.:,,:...-:0 i_iiiiiiiiiiiilii_iii!_ cO ::::2:2:::::::::: :,:,2.:-:0 i:_:_:_:i:i:i:::i_!iiiiig uJ_ _Oooo_ _.:,_ , • _. :i:!:i:i:i:!:_:_::::::::::_

":-:-:,:.:,:.:.:-:-:-:-:0 °° :':"'""/""-"0 %°°°IOo

:i:i_:i:!:!::::_

• - -:-:0:, i:i:=!:i:i:i b-l:_

___.: ..-...

::::O :1:: :-:0:-: LaJ_ ?:, n :.:-:..:.:,:.:. ¢-) U_ I f • _ ..... •o., : _

% .1-:-:-:.:,:.1.:,:£.:.1.:,: ::::::::::::::::::::::::: ,..,.....:,:.:.:.:.:.:,;.; 0 I

)6 '3_nlV_13dH31 _o '3_NL_3cIH3J 11.5 GAS BEARING PRESSURE DIFFERENTIAL DII-603 16 "_ 11.0 i!iiii_iiiii!ii_ii!ii!iiiiii{i!i - - I0.5 iiii!iiii;:_.iiii!!iiii:i::i::i::::ii::i15 ii!::::_ >>',:.:,:..... m I0.0 , II I I I _:"_:I"..... __ 14 _" 9.5

13.0 -19 12.5 _._....l.....L_J...... _..i l.....L 1....L....i...... t..... l...... _ ,--I-m--I-- PLATFORMINTERNAL AMBIENT PRESSURED12-603 ' i // ! _ -18 12.0 II.5 II.0 -16

-I 5 ._ 10.5 lO.O C_ -14 . 9.5 _= 9.0tI1"-...... ,...... !+...... +!...... -Ii,...... ---::!,...... _:...... -13

8.5 -12 8.0 ii_ii_-_!_i!_i_i_ii_i_i_!_i__ _ _-_6_C'Vt:i'GHT OPERATIN_RA_6E_:.i:f.:i_i-,i_!{iii -II 7.5 7.0 I0 6.5 2 4 6 8 I0 12 14 16 18 20 22 24 26 28 30 32 34 36 38 40 42 44 46 48 50 RANGE TIME, lO00 SECONDS

ii I I I I I I I I I I I I I 1:00:00 3:00:00 5:00:00 7:00:00 9:00:00 11:00:00 13:00:00 2:00:00 4:00:00 6:00:00 8:00:00 I0:00:00 12:00;00

RANGE TIME, IIOURS:MINUTES:SECONDS

Figure 18-7. Inertial Platform GN2 Pressures differential discrepancy is a function of a number of variables acting simultaneously with no single controlling factor. Although the gas bearing pressure regulator as a component fulfills its functional requirements, variables are introduced at a systems level which cause the pressure differential drift.

The occurrence of a slightly higher pressure differential on previous flights has resulted in no discernible effect on platform operation. Vendor testing of the inertial platform at pressure differentials up to 13.8 N/cm 2 (20 psid) have resulted in no degradation in platform performance.

An engineering change proposal is being considered to change the upper limit of the specification to an acceptable value, which includes the tolerance buildup of the system variables.

The GBS sphere pressure decay shown in Figure 18-8 was as expected. This was an indication of normal GN2 consumption.

18-9 22 -3200

-3000 2O 2800

18 2600 °. 2400 16 's 22Q0 _s 14 2000

% 1800 12 I m,- _" " ,./NOMINAL EXPECTED DECAY 1600 UO u_ I0 1400

8 1200 lO00

-800

6OO

4OO

2OO

0 4 8 12 16 20 24 28 32 36 40 44 48 RANGE TIME, I000 SECONDS

I:00:00 3;00:00 5:0Q;00 7:00:00 9:00:00 II:00:00 13:00:00 I I I I I I I I I I I I I 2:00:00 4:00;00 6:00:00 8:00:00 I0:00:00 12:00:00 RANGETIME, HOURS:MINUTES:SECONDS

Figure 18-8. GBS GN2 Sphere Pressure (D10-603)

18-I0 SECTION19

DATASYSTEMS

19.1 SUMMARY

All elements of the data system performed satisfactorily except for a problem with the Commandand Communications System (CCS) downlink during translunar coast.

Measurement performance was excellent, as evidenced by 99.9 percent reliability. This is the highest reliability attained on any Saturn V fl i ght.

Telemetry performance was nominal, with the exception of a minor calibration deviation. Very High Frequency (VHF) telemetry Radio Frequency (RF) propagation was generally good, though the usual problems due to flame effects and staging were experienced. Usable VHF data wer_ received to 17,800 seconds (04:56:40). Commandsystems RF performance for both the Secure Range Safety CommandSystems (SRSCS) and CCSwas nominal except for the CCSdownlink problem noted. Usable CCSdata were received to 35,214 seconds (09:46:54).

Goldstone (GDS) received CCSsignal carrier to 35,779 seconds (09:56:19). Good tracking data were received from the C-Band radar, with Patrick Air Force Base (PAFB) indicating final Loss of Signal (LOS) at 42,912 seconds (II :55:12).

The 75 ground engineering cameras provided good data during the launch. 19.2 VEHICLE MEASUREMENTEVALUATION

The AS-506 launch vehicle had 1370 measurements scheduled for flight. Eight measurements were waived prior to the start of the automatic countdown sequence leaving 1362 measurements active for flight. Of the waived measurements, five provided valid data during the flight.

Table 19-I presents a summary of measurement performance for the total vehicle and for each stage. Measurement performance was exceptionally good, as evidenced by 99.9 percent reliability, which is the highest attained on any Saturn V flight.

19-I Table 19-I. AS-506 Measurement Summary

MEASUREMENTS S-IC S-II S-IVB INSTRUMENT TOTAL CATEGORY STAGE STAGE STAGE UNIT VEHICLE

Scheduled 313 563 270 224 1370 Waived 3 4 1 0 8 Fai I ures 0 2 0 0 2 Partial Failures 8 0 0 0 8 Reli abi I i ty, I00.0 99.6 I00.0 I00.0 99.9 Percen t

Tables 19-2 and 19-3 tabulate by stage the waived measurements, totally failed and partially failed measurements. None of the listed failures had any significant impact on postflight evaluation. 19.3 AIRBORNETELEMETRYSYSTEMS

Performance of the eight VHF telemetry links was generally satisfactory with the minor exceptions noted. A brief performance summary of these links is shown in Table 19-4.

There was a variation of approximately 17 counts in the I00 percent level of the inflight calibrations for the DP-I telemetry link. This is equiv- alent to 85 millivolts as compared to 41 millivolts in the specifications. This type of variation is present in all other calibration levels to a lesser degree. The data indicate the variation is from the Model 301 or the Model 270 multiplexers and not the 5-volt measuring supply. This problem, which also occurred on AS-505, is being investigated.

Data degradation and dropouts were experienced at various times during boost as on previous flights due to attenuation of RF transmission, at these times, as discussed in paragraph 19.4.1.

Usable VHF telemetry data were received to 17,800 seconds (04:56:40) at Guaymas (GYM).

Performance of the CCS telemetry was generally satisfactory except for the period during translunar coast from 27,128 seconds (07:32:08) to 35,779 seconds (09:56:19). This problem is discussed in detail in paragraph 19.4.3.2. Usable CCSdata were received at GDSto 35,214 seconds (09:46:54).

19-2 Table 19-2. AS-506 Flight Measurements Waived Prior to Launch

MEASUREMENT NUMBER MEASUREMENT TITLE NATURE OF FAILURE REMARKS

S-IC STAGE

D004-I02 Pressure, Fuel Pump Transducer offset and not KSC Waiver I-B-506-3. Meas- Inlet l responsive to pressure urement provided valid data throughout powered flight Dllg-102 Pressure Differential, Transducer fai Iure KSC Waiver I-B-506-3 Engine Gimbal System Filter Manifold DI19-I03 Pressure Differential, Transducer fai Iure KSC Waiver I-B-506-3 Engine Gimbal System Filter Manifold

S-II STAGE

! C_ D008-201 El LOX Turbine Inlet Transducer drift Flight data usable Pressure DI04-201 Engine Hydraulic Noisy transducer Flight data usable Reservoir Pressure DI04-202 Engine Hydraulic Noisy transducer Flight data usable Reservoir Pressure

D_04-203 Engine Hydraulic Noisy transducer Flight data usable Reservoir Pressure

S-IVB STAGE

C0005-405 Temp. - Cold He Measurement failed off-scale It is suspected that a Sphere No. 3 Gas low during LH2 loading of resistance dependent upon CDDT. temperature was shunting the temperature sensor. This causes a lower than calibrated resistance of the probe to be seen by the bridge. Table 19-3. AS-506 Measurement Malfunctions

TIME OF MEASUREMENT MEASUREMENT SATISFACTORY NATUREoF FAILURE FAILURE REMARKS NUMBER TI TLE OPERATION (RANGE TIME) TOTAL MEASUREMENTFAILURES, S-II STAGE C003-205 E5 Fuel Turbine Transducer opened S-II ESC 0 second Inlet Temp. F001-204 E4 Main Fuel Flow No data pulses Prior to launch 0 second during engine burn PARTIAL MEASUREMENTFAILURES, S-IC STAGE

C003-I Ol Temperature, No data from 0 to 77 0 second 85 seconds Probable cable Turbine Manifold seconds problem DO07-101 Pressure, Fuel Data decreases after 135 seconds 135 seconds Probable trans- Pump Discharge 2 135 seconds ducer malfunction D007-I02 Pressure, Fuel Data approximately 105 seconds 148 seconds Probable trans- I Pump Discharge 2 I00 psi low from ducer malfunction 105 to 120 seconds D007-I05 Pressure, Fuel Data approximately 85 seconds 148 seconds Probable trans- Pump Discharge 2 I00 psi low from ducer malfunction 85 to I00 seconds

DO16-I 04 Pressure, Engine Data approximately 95 seconds 122 seconds Probable trans- Gimbal System I00 psi high from ducer malfunction Supply, Engine 4 95 to 135 seconds DI18-I04 Pressure, Engine Data erratic subse- 140 seconds 140 seconds Probable trans- Gimbal System quent to 140 seconds ducer malfunction Return, Yaw Actuator

D144-I19 Pressure, Helium Data read low from I00 seconds 132 seconds Cause unknown Storage Tank I00 to 130 seconds

K085-120 LOX Tank Vent Data noisy and Ignition 27 seconds Probable cable Valve, Closed erratic from igni- problem tion to 135 seconds Table 19-4. AS-506 Launch Vehicle Telemetry Links

FREQUENCY MODULATION STAGE FLIGHT PERIOD LINK MHz (RANGETIME, SEC) PERFORMANCESUMMARY

AF-I 256.2 FM/FM S-IC 0-410 Satisfactory AP-I 244.3 PCM/FM S-IC 0-410 Data Dropouts Range Time (sec) Duration (sec) 162.3 l.O 165.5 1.7

BF-I 241.5 FM/FM S-II 0-772 Satisfactory BF-2 234.0 FM/FM S-II 0-772 Data Dropouts BP-I 248.6 PCM/FM S-ll 0-772 Range Time (sec) Duration (sec) 163.0 2.5 192.3 3.0

I CP-I 258.5 PCM/FM S-IVB Flight Duration Satisfactory Data Dropouts Range Time (sec) Duration (sec) 162.4 1.0

DF-I 250.7 FM/FM IU Flight Duration Satisfactory except for DP-I DP-I 245.3 PCM/FM IU Flight Duration calibration. DP-IB 2282.5 PCM/FM IU 0-35,779 Data Dropouts Range Time (sec) Duration (sec) 162.9 (VHF) 2.1

162.5_ 7.0 193.5 8.0 17,470 See 19.4.3.2 27,128 'DP-IB See 19.4.3.2 30,264 only See 19.4.3.2 34,020 See 19.4.3.2 35,214 j See 19.4.3.2 19.4 RF SYSTEMS EVALUATION

19.4.1 Telemetry System RF Propagation Evaluation

The performance of the eight VHF telemetry links was excellent and generally agreed with predictions. VHF telemetry links AF-2, AF-3, AS-I, AS-2, BF-3, BS-I, BS-2, CF-I and CS-I were deleted on AS-506.

Moderate to severe signal attenuation was experienced at various times during the boost due to main flame effects, S-IC/S-II and S-II/S-IVB staging, S-II ignition and S-II second plane separation. Magnitude of these effects was comparable to that experienced on previous flights. At S-IC/S-II staging, signal strength on all VHF telemetry links and on the CCS downlink dropped to threshold for approximately 2 and 7 seconds, respectively. Signal degradation due to S-II ignition and S-II flame effects was sufficient to cause loss of VHF telemetry data on the S-IC and S-II stages. CCS and S-II VHF data were lost during S-II second plane separation. In addition, there were intervals during the launch phase where some data were so degraded as to be unusable. Loss of these data, however, posed no problem since losses were of such short duration as to have little or no impact on flight analysis.

The performance of the S-IVB and IU telemetry systems was nominal during orbit, second burn and final coast, except for the CCS problem discussed in paragraph 19.4.3.2.

GYM reported VHF LOS at 17,800 seconds (04:56:40) and GDS reported CCS LOS at 35,779 seconds (09:56:19).

A summary of available VHF telemetry coverage showing Acquisition of Signal (AOS) and LOS for each station is shown in Figure 19-I.

19.4.2 Tracking Systems RF Propagation Evaluation

Analysis of data received to date indicates that the C-Band radar functioned satisfactorily during this flight, although several ground stations experienced some tracking problems.

The only problems reported during launch occurred at Cape Kennedy (CNV), Merritt Island Launch Area (MILA), and Grand Turk Island (GTK). All three stations lost track due to balance point shifts (erroneous pointing information caused by a sudden vehicle antenna null or a distorted beacon return). CNV and MILA went off track momentarily at I00 and 395 seconds, respectively. GTK had dropouts due to balance point shifts at 241 seconds (momentarily), from 535 to 538, from 555 to 570, from 572 to 580, from 594 to 599 and from 606 to 614 seconds. The highest elevation angle encountered by GTK during this period was 3 degrees. MILA went off track from 440 to 480 seconds due to interference from an electrical storm. Bermuda (BDA) did not report any problems during launch.

19-6 BEGIN S-IVB RESTART PREPARATIONS S-IVB SECOND IGNITION PARKING ORBIT INSERTION TRANSLUNAR INJECTION

GYM TAN CYI VAN BDA

ClF

I t I i I i t I f 0 600 1200 1800 2400 3000 3600 4200 4800 5400

RANGE TIME, SECONDS ST7 I IY I I I I I I I I 00:00:00 00:10:00 00:20:00 00:30:00 00:40:00 00:50:00 01:00:00 01:10:00 01:20:00 01:30:00

RANGE TIME, HOURS:MINUTES:SECONDS

MER I CRO

I TAN

CYI VAN TEX

I I I I I I I I I 5400 6000 6600 7200 7800 8400 9000 9600 1O, 200 10,800 RANGE TIME, SECONDS

I I I I , , i x_7 ,V,V , 01:30:00 01:40:00 01:50:00 02:00:00 02:10:00 02:20:00 02:30:00 02:40:00 02:50:00 03:00:00

RANGE TIME, HOURS:MINUTES:SECONDS

I TEX I HAW

GYM

I I I I I I I t I I 7200 10,800 14,400 18,000 21,600 25,200

RANGE TIME, SECONDS

t I I I I I I I I I I 02:00:00 03:00:00 04:00:00 05:00:00 06:00:00 07:00:00

RANGE TIME, HOURS:MINUTES:SECONDS

Figure 19-I. VHF Telemetry Coverage Summary

19-7 Problems experienced during earth orbit included tracking on a sidelobe by Vanguard (VAN) (revolution 2) and a phasing problem experienced by GTK (revolution 2). This type phasing problem is experienced when a ground station receives two closely spaced beacon returns; one generated as a result of its own interrogation and one resulting from the interrogation of the beacon by another ground station.

GTK lost track during translunar coast from 27,126 seconds (07:32:06) to 29,260 seconds (08:07:40) when attempting to phase away from the beacon return pulse of another ground station. PAFB indicated final LOS at 42,912 seconds (11:55:12).

A summary of available C-Band radar coverage showing AOSand LOS for each station is shown in Figure 19-2.

There is no mandatory tracking requirement of the CCS; however, the CCS transponder has turnaround ranging capabilities and provided a backup to the Commandand Service Module (CSM) transponder used for tracking in case of failure or desire for a cross check. Since the same transponder is used for all CCSfunctions, discussion of the tracking performance of this system is included in the general discussion of the CCSRF evaluation.

19.4.3 CommandSystems RF Evaluation

19.4.3.1 SecL_re Ranqe Safety Command System. VHF telemetry measurements received by the ground stations from the S-IC, S-II and S-IVB stages indicated that the SRSCS RF subsystems functioned properly. CNV and BDA were the command stations used for this flight. The carrier signal at CNV was turned off at approximately 400 seconds. At BDA the carrier was turned on at approximately 375 seconds and turned off at approximately 750 seconds. A momentary dropout occurred at approximately 120 seconds when the command station switched transmitting antennas.

19.4.3.2 Command and Communications System. Available data indicated satisfactory CCS performance during boost and parking orbit with minor exceptions. Uplink and downlink dropouts occurred during S-IC/S-II staging and at S-II second plane separation. Dropouts at these times are expected. Performance during second burn and translunar injection was nominal.

Signal fluctuations were noted at HAW, GBM, GDS, and GYM from about II,I00 seconds (03:05:00) to II,340 seconds (03:09:00) when the CSM was maneuvered to an inertial attitude. This inertial attitude was maintained during CSM separation, docking and Lunar Module (LM) ejection.

HAW lost track during translunar coast from II,756 seconds (03:15:56) to 18,516 seconds (05:08:36) when the vehicle disappeared over the horizon.

19-8 BEGIN S-IVB RESTART PREPARATIONS S-IVB SECOND IGNITION TRANSLUNARPARKING ORBITINJECTIONINSERTION VAN FPS-I 6

BDA FPQ-6 BDA FPS-I 6 PAFB FPQ-6 MILA TPQ-18 CNV FPS-16 CRO FPQ-6

i i I I I I I I 600 1200 18OO 2400 3000 3600 4200 4800 5400

RANGE _IME, SECONDS

I _ , I I I I I I I I 00:00:00 00:I0:00 00:20:00 00:30:00 00:40:00 00:50:00 01:00:00 01:I0:00 01:20:00 01:30:00

RANGE TIME, HOURS:MINUTES:SECONDS

VAN FPS-16 BDA FPQ-6 PAFB FPQ-6 MILA TPQ-18 CRO FPQ-6

t , I I I , I I I I I 54OO 6000 6600 7200 7800 8400 9000 9600 10,200 10,800

RANGE TIME, SECONDS

I I I I , , , V , 01:30:00 01:40:00 01:50:00 02:00:00 02:10:00 02:20:00 02:30:00 02:40:00 02:50:00 03:00:00

RANGE TIME, HOURS:MINUTES:SECONDS

HAW FPS-16 ASC TPQ-18 ANT FPQ-6 BDA FPQ-6 PAFB FPQ-6

I I t I I I I I [ I 7200 10,800 14,400 18,000 2_,600 25,200 28,800 32,400 36,000 39,600 43,200

RANGE TIME, SECONDS

I I I ! ! I I i I t I 02:00:00 03:00:00 04:00:00 05:00:00 06:00:00 07:00:00 08:00:00 09:00:00 I0:00:00 11:00:00 12:00:00

RANGE TIME, HOURS:MINUTES:SECONDS

Figure 19-2. C-Band Radar Coverage Summary

19-9 A ground commandwas transmitted at 17,466.6 seconds (04:51:06.6) to initiate Time Base 8. The vehicle was placed in a slingshot attitude and the LOX dump followed. These events produced signal strength fluctuations from 17,470 seconds (04:51:10) to 19,060 seconds (05:17:40.) at all stations tracking the CCS. The most severe fluctuations were experienced at GBMand resulted in 25 dropouts during this time period. These signal fluctuations were smooth and are believed to have been caused by changing vehicle antenna gains as the look angles to the ground stations varied with the changes in vehicle attitude (referenced to the ground station).

A sharp drop in downlink CCSsignal was noted at HAW, GBM,GDSand GYMat 27,128 seconds (07:32:08). The onboard antenna system, which had been on the low gain since 19,034 seconds (05:17:14) was switched to the high gain mode at 27,368 seconds (07:36:08) to improve signal quality. Signal strength picked up and was maintained at a high level until 30,264 seconds (08:24:24) at which time the signal level again dropped. In an attempt to improve signal quality the CCSRF was switched OFF/ONtwo times and the CCSantennas were switched several times. However, signal level fluctuated intermittently at low levels until LOS at 35,779 (09:56:19). Figure 19-3 shows the fluctuations in signal level experienced at the HAWsite. The GDSwing station experienced similar fluctuations at corresponding times as shown in Figure 19-4. CCS RF OMNI

CCS RF OFF HIGH GAIN 0 OMNI CCS RF ON i

i -113

E _Q

-123 -103 . .

Hyll .. , -lS3 i, 27,000 28,000 29,000 30,000 31,000 32,000 33,000 34,000 35,000 36,000 RANGE TIME, SECONDS

I I I I I I 7:30:00 8:00:00 8:30:00 9:00:00 9:30:00 I0:00:00 RANGE TIME, HOURS:MINUTES:SECONDS Figure 19-3. CCS Signal Strength Fluctuations at Hawaii

19-I0 CCS RF OMNI ON _LOW GAIN HIGH GAIN OFF _CCS RF OFF OMNI CCS RF ON

-I O0

-150

-160 __ 27,000 28,000 29,000 30,000 31,000 32,000 33,000 34,000 35,000 36,000 RANGE TIME, SECONDS

I I I I I I 7:30:00 8:00:00 8:30:00 9:00:00 9:30:00 I0:00:00 RANGE TIME, HOURS:MINUTES:SECONDS Figure 19-4. CCS Signal Strength Fluctuations at GDS Wing Station

The above indicates that the problem was present on low gain, high gain and omni antenna; therefore, it is concluded that the drop in signal level was caused by a malfunction of the CCS coaxial switch. On AS-505, a similar problem in the CCS antenna system occurred only while transmitting on the high gain or low gain antenna.

Test performed in IBM Report Number 69-223-0007 also concluded that the CCS coaxial switch (the only electromechanical component which is common to all CCS antennas) caused the failure. The general characteristics of the CCS operation, as observed on AS-505 and AS-506, was duplicated by a simulated leak in the hermetically sealed portion of the coaxial switch case. In addition, engineering tests have demonstrated that the coaxial switch will leak following vibration levels seen on AS-505 and AS-506.

19-11 Directional antenna tests did not duplicate the failure. Power amplifier tests showed a leak in the power amplifier would cause a total failure; this results in total loss of CCSdownlink with no possible recovery.

Prior to any observed deficiencies in the flight operation of the CCS, incorporation of a new design coaxial switch was programed for AS-507 and subsequent vehicles. The new switch exhibits none of the general deficiencies of the earlier components and has shown no susceptibility to failure in simulated leak tests or at vibration levels in excess of the AS-505 or AS-506 vibration levels.

A summary of CCScoverage showing AOS and LOS for each station is shown in Figure 19-5.

19.5 OPTICAL INSTRUMENTATION

In general, ground camera coverage was very good. Seventy-five items were received from Kennedy Space Center (KSC) and evaluated. One camera jammed before acquiring requested data. Two cameras had bad tracking items, one camera had its field of view misoriented and one camera had no run. As a result of the 5 failures listed above, system efficiency was 94 percent. All Launch Umbilical Tower (LUT) cameras had erratic timing; therefore, all timing data were interpolated. Personnel at KSChave traced the timing problem to a loose connector at the base of the Launch Control Center (LCC).

19-12 BEGIN S-IVB RESTART PREPARATIONS S-IVB SECOND IGNITION PARKING ORBIT INSERTION TRANSLUNAR INJECTION

m CYI n VAN BDA IN HSK GYM MI LA CRO n GDS WING STA. I I I I I I I I I - I, 0 600 1200 1800 2400 3000 3600 4200 4800 5400 6000

RANGE TIME, SECONDS

I n_ 7 I I I I I i I I 00:00:00 00:10:00 00:20:00 00:30:00 00:40:00 00:50:00 01:00:00 01:I0:00 01:20:00 01:30:00

RANGE TIME, HOURS:MINUTES:SECONDS

m _DS m VAN m BDA

m GBM ml_ MILA I i I I I I I I I 5400 6000 6600 7200 7800 8400 9000 9600 10,200 10,800 RANGE TIME, SECONDS

I I I I I I I _7 i_ 7 _ I / 01:30:00 01:40:00 01:50:00 02:00:00 02:10:00 02:20:00 02:30:00 02:40:00 02:50:00 03:00:00 RANGE TIME, HOURS:MINUTES:SECONDS GYM GDS WING STA. GDS n HAW , , HAW

II MER HAW TRACKED CSM CCS II CRO CYI GBM

II_ MILA

I I I I I I I I I 7200 10,800 14,400 18,000 21,600 25,200 28,800 32,400 36,000 39,600 RANGE TIME, SECONDS

I I I I I I I I I I 02:00:00 03:00:00 04:00:00 05:00:00 06:00:00 07:00:00 08:00:00 09:00:00 10:00:00 II:00:00

RANGE TIME, HOURS:MINUTES:SECONDS

Figure 19-5, ,CCS Coverage Summary

19-13/19-14

SECTION20

MASSCHARACTERISTICS

20.1 SUMMARY

Postflight analysis indicates that total vehicle mass was within 0.50 percent of the prediction from ground ignition through S-IVB stage final shutdown. This very small deviation signifies that the initial propellant loads and propellant utilization throughout vehicle operation were close to predicted. 20.2 MASSEVALUATION

Postflight mass characteristics are compared with the final predicted mass characteristics (MSFCMemorandumS&E-ASTN-SAE-69-M-70) and the final operational trajectory (MSFCMemorandumS&E-AERO-FMT-138-69).

The postflight mass characteristics were determined from an analysis of all available actual and reconstructed data from S-IC stage ignition through S-IVB stage second burn cutoff. Dry weights of the launch vehicle were based on actual stage weighings and evaluation of the weight and balance log books (MSFCForm 998). Propellant loading and utilization was evaluated from propulsion system performance reconstructions. Spacecraft data were obtained from the Manned Spacecraft Center (MSC).

Deviations from predicted in dry weights of the inert stages and the loaded spacecraft were all less than 0.75 percent which was well within the 3-sigma deviation limit.

During S-IC powered flight, mass of the total vehicle was determined to be 2906 kilograms (6407 Ibm) or 0.09 percent lower than predicted at ignition, and 1366 kilograms (3011 Ibm) or 0.16 percent lower at S-IC/ S-II separation. These small deviations are attributed to less than predicted S-IC propellant load, S-IC dry stage mass, and mass of the upper staging. S-IC burn phase total vehicle mass is shown in Tables 20-I and 20-2.

During S-II burn phase, the total vehicle mass varied from 898 kilograms (1981 Ibm) or 0.13 percent lower than predicted at start commandto 875 kilograms (1930 Ibm) or 0.42 percent higher than predicted at S-II/S-IVB separation. Most of the initial deviation may be attributed to a less than predicted S-II propellant loading, and the deviation at separation

20-I was due mainly to higher than predicted S-II propellant residuals. Total vehicle mass for the S-II burn phase is shown in Tables 20-3 and 20-4.

Total vehicle mass during both S-IVB burn phases, as shown in Tables 20-5 through 20-8, was within 0.45 percent of prediction. A deviation of 143 kilograms (317 Ibm) or 0.09 percent at first start command was due mainly to a slight excess of S-IVB propellants. Lower than predicted propellant residuals at end of first burn resulted in a 607 kilogram (1340 Ibm) or 0.44 percent deviation. Total vehicle mass at spacecraft separation was 832 kilograms (1834 Ibm) or 4.62 percent less than predicted.

A summary of mass utilization and loss, actual and predicted, from S-IC stage ignition through completion of S-IVB second burn is presented in Table 20-9. A comparison of actual and predicted mass, center of gravity, and moment of inertia is shown in Table 20-10.

20-2 Table 20-I. Total Vehicle Mass - S-IC Burn Phase - Kilograms

RRBUNn IGNITION HOLDr)OWN CENTFR OUTBOARD _-IC/S-I I EVENTS ARM RELFASE ENGINE CUTOFF ENGINE CUTOFF SEPARATION

PRE_ AC [ PRED ACT PRFn AC T PRFn ACT PREO ACT

RANGE TI ME--_E C -8 ,. q, 0 -6o40 .30 °30 135.28 135.20 161,.08 161.63 161.,80 162.30

S-IC _TAGE DRY 130975. 130522. 130975. 130522. I _0975. 130522. 130975. 130522. 130975. 130522. LOX IN TANK 1579518. 1578371. 1558229. 1558728o 190236. 195782. 1399. 1280,, 931. 935. LOX BELOW TANK 2]000. 21108. 21737. 2 1868. 21720. 21851. 18778. 18761 . 14883. 14717. LOX ULLAGE GAS 187. 169. 207. 235. 2587. 28{19. 31160. 3611. 3088. 3616. RPI IN TANK ¢52551. 642018. 632357. 631857. 91469. cl 3r) 78. 8396. 8008. 7305. 8759. RP! BELOW TANK 5313. 4301. 5996,, 5983. 5596. 5983. 5958. 5958. 5958. 5956. RPl ULLAGE GAS X5. 73. 35. 76. 21 1. 226. 250. 259. 241. 250,, N2 PURGE GAS X6. 36. 36. 38. 20. 70. 20. 20. 20. 20. HELIUM IN BOTTLE 289. 289. 289. 28&. 113. 13G. 83. 112. 83. 112. FROST 6 X5. 835. 635. 635. 35 O. 3 _O. 3_ O. :3 40. 34 O. 350. RETROMOTOR 1027. 1027. ID27. 1027. 1027. 1027. Z027. 1027. 1027. 1027. r%) PROP C) OTHER 2 ]q. 239. 239. 2_q. 239. '2 39. 235. 2 39. 239. 2 39. I r.o TOTAL S-IC STAGE 2280695. 2278688. 2251801. 22X9381. 445932. 450512, 16851_. 168015. 165857. 16_381.

TOTAL S-ICIS-II IS 5200. 5208. 5200. 5206. 5200. 5206. 5?00. 5208. 5166. 5173. TOTAL S-II STAGE 481"003. 47998q. 481003. 575964. qR07q_5. 4797(16. 480755. 47971]6. 480745. 479708. TOT S-II/S-IVB IS 3665. 3883. 3665. 3663. 3665. 3863. 3665. 3663. 3665. 3863. TOTAL S-IVR STAGE 118911. 119119. 118911. 119119. 118820. 1191129. 118820. 119029. 118820. 119029. TOTAL INSTRU UNIT I953. 1939. 1953. 1939. 1953. 1539. 1953. 1939. 1953. 1939. TOTAL SPACECRAFT 45754. _9735. 49794. _9735. 49794. q 9735. 597911. 49735. 49795. 597 35.

TOTAL UPPER STAGE 660526. 859826. 660526. 859826. 860177. 655277. 660177. 659277. 68015, q- 659245.

TOTAL VEHICLE 2941221. 2938315. 2902328. 2899008. 11115110. 1110189. 828692. R27292. 824991. 823625. "£8LSIBI "h6L818! "_98£Z8! "h569_8[ °LhbLtphZ "0S£9_1_Z °t/lZl6£9 °LES8619 "SLBLLW9 "_uZhSh9 3]:JIH]A ]VIOl

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ssEH spunod - asEqd uJn8 3I-S - ssEH ato._qaA LE:_Ol "_-0_ atqEl Table 20-3. Total Vehicle Mass - S-II Burn Phase - Kilograms

S-IC IGNITION S-I I S-II S-II S-I I/R-IV_ EVENTS IGNITION MAT NSTAGE ENGINE CUTOFF SEPARATION

PRED ACT PRED ACT PRFD ACT PRED ACT PRED AC T

RANGE TI ME--SEC -6.qD -G. (I0 ]63.54 ]Gq.OD IG5.50 I66.20 551.72 548.22 552. qO 549.00

S-IC/S-II IS SMALL BIq. 614. S-IC/S-II IS LARGF 39&9. 39R2. 3969. 398?° 3969. 3982. S-IC/S-II IS PROP 617. BID. 3]3. 309. 0. 0.

TOTAL S-IC/S-II IS 5200. S2O&. 4281. 4291. 39S9. 3982. o S-TI STAGE DRY 36250. 36158. 3625n. 36158. 3625D. XGlS8. 3625n. 38158. 36250. _lSR. ! o'I LOX IN TANK 371_72. 370778. 371672. X7{]77R. 371221]. 370325. 657. 816. Sqq. 730. LOX BELOW TANK 7_7. 737. 737. 737. 800, 8110. 787. 787. 787. 787. LOX ULLAGE GAS 188. 1R8. IRR. 18R. 190. 191. 2"_37. 2335. 23{10. 2335. LH? IN TANK 71668. 71615. 71660. 71608. 71449. 71396. I966. 2572. 1916. 2531. LH? BELOW TANK I05. I05. 112. liP. 128. 128. 123. 123. 123. 123. LH? ULLAGE GAS 77. 77. 77. 77, 77. 78. 704. 715. 7oq. 735. INSULATION PURGE 54. sq. FROST 70li. 2114. START TANK GAS 14. ]4. ]q. ]q. 2. 2. 2. 2. 2. 2. OTHER 34. 34. 34. 34. 3 q. 34. "_q. 34 . 3 4. 34 .

TOTAL S-II STAGE 4A]0113. 479964. 480745. 4797flG. 4R0]51. 4791 12. 428&2o 43564. _2702. 4343&.

TOT S-II/S-IVB IS ]665. 3663. 3665. 366.X. 3R65. 3_63. 3565. 3663. 3465. 3&63. TOTAL S-IVB STAGE llBgll, llqllg, llBB20. 119029. 118820. llqr179. 118820. 119029. 118818. 119026. TOTAL INSTRU UNIT I951. 1939. 1953. 1939. 1953. 1939. 1953. 1439. 1953. 1939. TOTAL SPACECRAFT 497_q. 49735. 49794. 49735. 4979 4. 49735. 4574 3o _15693 . 4574 3. W5693.

TOTAL UPPER STAGE ]74324. |74456. 174233. 174365. ]74233. 174365. 170182. 170324. 170180. 170322.

TOTAL VEHICLE 660526. 659626. 659259. 658363. 658 _53. G57fi 59. 7 tXO_q. 2l 3888. 212882. Z| 3757. " bSZIL,; "hZ£GBh "ZbSILh "ZB969b "OSb6hb[

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ss_H spunod - as_qd uanEl II-S - ss_IAl aLD.LqaA L_Ol "17-0_ aLq_l Table 20-5. Total Vehicle Mass - S-IVB First Burn Phase - Kilograms

S-IC IGNITION S-IVB S-IVB S-IVB 5-IVB EVENTS IGNITION MAINSTAGE ENGINE CUTOFF END DECAY

PREO ACT PRED ACT PRED AC T PRF(] ACT PRED ACT

RANGE TI MF--SE C -6.qO -6.40 555.,78 552.7n 55A. 2D 554.7n 699o49 699.34 699.,68 699.54 ___ S-IVB STAGE DRY 11340. 11273,. 11317. 1125(1. II317. II?5D. 11255. II189. 11255,, 11189. LOX IN TANK 86q34. 87149. 86934. 87149. 8R773. 86993,, 61359. 61120. 61327. 61052. o LOX BELOW TANK IGG. 166. IGG. l&&. 180. 180,, 180. 180. 180., 180. I LOX ULLAGE GAS 17. 16. 17. IG. 22., 17. 105,, 67. lOS. _7. LH2 IN TANK 19709. 19758. 197(15. 13731. 19644. 19708. 14530° 14]69° 14516. 14]56. LH? BELOW TANK 27,, 22. 26. 26. 26. 26. 26. 2E. 26. 26. LH2 ULLAGE GAS 2n. 19,, 20,, 19. 20., 20. 65. 52. E6. 52. ULLAGE MOTOR PROP 54. 5_. IO. In. I. I. I. 1. I. 1. APS PROPELLANT 28_. 298. 2RG. 29R. 286o 298. _R5° 297,, 285° 297. HELIUM IN BOTTLES 20_1. 200. 20D. 200. 199. 200° ITR° 176. 178,. 17E. START TANK GAK 2. 2. _. 2° {_. D. _° 3. 3. 3. FROST ] __G. 13&. 45. 45. 45,, 45. 45,, 45. 45. 45. OTHER 25. 25. 25. 25. 25. 25. 75. 25. 25. 25.

TOTAL S-IVR SIAGF 118911. llqllq. 118754. 118939. llRSq4. I187G4. 88058° 87550. _8{)13° 87469.

TOTAL INSTRtJ UNIT 1953. 1939. 1953. 193'q. 1953. lq]9. 195_. 1939. 195_° 19]9.

TOTAL SPACECRAFT 45743. 45693. 45743. qs&q3. 45743° 45693° 45743. 45693,, NS7q]. 45693.,

TOTAL UPPER STAGE 49697. 47&32. 47_97. 47&32. 47697. 47_32. 47697. 47632° 476970 _7632.

z,-- TOTAL VEHICLE lG&EnB. 166751,, 1REqSt]. 166571. ]AEPqD. 16_X96. 135755., I._51 82. l ]5709° 13_102. Table 20-6. Total Vehicle Mass - S-IVB First Burn Phase - Pounds Mass

%-It. IGNITION S-IVR S -°[VR S-IVR S-IVR FVENTS IGNI [ION MA [NqT AGE FNGINE CUTOFF END DECAY

pRErl AC T PP ED AC T PR FI_ AC T PRFD AC T PRED AC T

RANGE TI ME--SEC -6.40 -Go =40 555.70 552.211 SSR. 20 554o7n 699,.49 &99.34 649.68 699.54

S-IVB STAGE DRY 2500II. 2=4852. 24944. 24801. 24949. 24801. 24_14., 24667,, 24814. 24667. LOX IN TANK Iq1656. 192130. IqIG5R. 197130. ]91.X02. 191767. 135773. 134747. 135203. 134597. LOX' BELOW TANK 367. 367. 367° 367. 397. 347. x97. 397. 397. 397,, LOX ULLAGE GAS T_R. 36. 36. 36. 49° 38° 231,, 147,. 232,,. 147° LH? IN TANK 43452. 435G0. 434w2. 43499. 43X18. 43449. 370X3. 31678. 32002- 31649. LH2 BELOW TANK _R. 48. 5B. 5R. 54. 58,, .58. 58. 58,, 56. LH9 ULLAGE GAS qX° 41. 43. 42. 44. 44. 144. llq. 145° 115. ULLAGE MOTOR PROP 116. liB. 27. 22. 0. O° tl. O° O. O. APS PROPELLANT G30. 65R. 630. 65R. _30. 65_. R2B. 655. G28. 655. HELIUM IN BOTTLES _ql. 442. 441. 442° 4_9. qql. xgx. XBq. 392. 389. r_ START TANK GA_ 5° 5. 5° 5,, 1. 1. 7. 7. 7° 7. c) FROST 306. 300. 100. 1DO. 100. TriO. 1011. 100. 100. 100. I OTHE R 56. 56. 56. 56 . 5 6. 56. 5 G. 56. 5 6. 56 . Co

TOTAL S-IVR STAGE 262154. 262613. 7618n7° 267216. ?G1344. 761830. 194135. 1911015. 194r135° 192837.

TOTAL INSTRU UNIT 4306. 4275. 4306. 4275. 4306. 4275° 4306,, 4275. 4306° 4275.

TOTAL SPACECRAFT I008_7. 100736° 100847° 100736. I Q0847. IOQ736o I00_47. I00736. I00847. 100736.

.,,2---- TOTAL UPPER STAGF I05153. I05011. 105153. IOSOll° 1II5153. I05011. I_fi153° IO_OIIo 105153. lOSOll.

TOTAL VEHICLE 367307. 367624° 366960. 3672P7. 3_6497. 366641. 299288. 298026° 299188. 797848. Table 20-7. Total Vehicle Mass - S-IVB Second Burn Phase - Kilograms

SPACECRAFT S-IVP S-TVB S-I VR _-IV9 EVENTS IGNITION MAINSTAGF ENGINE CUTOFF END DECAY SEPARATION

PRED ACT PRED ACT PRED ACT PREO ACT PRED ACT

RANGE TIME--SEC 985S.S0 985G.20 9858.00 9858.70 10704.0G 10203.07 10204°27 I0203.27 15004.40 15_73o00

S-IVB STAGE DRY 11255. 11189. 11255. II189. 11255. I1189. 11755. 11189. 11255. 11189. LOX IN TANK Gl2qO. 60985. gi074. G0857. 2191., 2.X08. 21GO° 2247° 2160. 2224. 0 I LOX BELOW TANK IgG. IGG. 180. 180. 180° 180. 180. 180,, 1G6. IG6. ',.o LOX ULLAGE GAS 170. 12G. 174. 128. 280. ?I!5° 280. 20_. 280. 124. LH? IN TANK 13257. 13275. 13192,, 13224. 900. 944. 886,, 932. 886. 391° LH2 BELOW TANK 26° 2G. 2G. 2G° 26° ?G,. 2G. 26. 22. 22,, LH? ULLAGE GAS lqg. IG?. IBT. 170. 331. 2RG. 331. 28G° 331. 15&. ULLAGE MOTOR PROP O. O. n. O. O. O. O. O. O. O. APS PROPELLANT 1R1;. 24G. 183. 246. 179. :_37° 179° 237. 144. 722. HELIUM IN BOTTLES I_5. IGO. 1W5. 159. 83. 1.98. 83. 108,, 8_X. 17. START TANK GAS 2. 2. O. O. 3. 3. 3. _. 3. O. FROST 4_ . 45. 45. 45 . 45. _5 . 4 5. 45. 4 5. 45. OTHER 25. 25. 2_. 25. 25. 25,, 25. 25 . 2 5. 25.

TOTAL S-IVR %TAG_ 89711. 8G419° 86_97. 8G251. 15500° 15557° 15454° 15983. 15401. 14583.

TOTAL INSTRU UNIT 1953. 1939. 1953,, 1939. 1953,. I939° 1953,. 1939° 1953. 1939.

TOTAL SPACECRAFT 457W3. 45G93° W5743. 45693. 4574 3. 4_93. 45743. 4S693. 626. G2g°

.... ,r TOTAL uPPER STAGE 47_a7° 971=::;32. 47G97. _7632. 47G97. 47G32° _7697. 47G32. 2579. 25G5.

TOTAL VEHICLE 134_f_8. 1 3qo_G. 134194. 133883° 63 1,9 6. £3189,, G3151. 6._11G. 17980. 171q8° iI

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• • • " • i¸!iiii_: • _ Table 20-9, Flight Sequence Mass Summary

PPE D] CTFD AC TU AL MASS HISTORY KG LBM KG LBM

S-IC STAGE, TOTAL 2280695° 5n? 8071. 2278688. 50236q8° S-ICI_-II INTFRSTAGE,TO_AL 5;_00. I 1qG3. 5706. ] 1q77. S-II STAGE, TOTAL q81003° 10__0q31. q799G_W° | 058]qO. S-II/_TIVB INTERSTAGE 3665° 8DRI. 3663. 8076° S-IVB STAGEr TOTAL 118911. 26215q. 11911q. 262613. INSTRUMENT UNIT 1q53° q3OGo 193q. q275. SPACECRAFI INCLUDING LFS q979q. 109777. q9735, lfl 9F_q6°

IST FLT STG AT IGN 279q 127. 1 ° 6qBq2782. 2938315. 6q77875. S-IC THRUST BUILDUP -38893. -857q5. --3q307. --866_7.

IST FLT GIG HOLDDWN ARM REL 7gr12328° 63985x7. 2899008 ° 63912718. S-IC FROST -295 . -KSO . --295 . -650. S-IC MAINSTAGE PPOPFLL_A, NT -2071872. -_567_97. -2069957° --qS63qTqw _-IC N2 PURGF - 17. -37. -17. -37. S-IC INBD ENGINE T.D. PROP -q 17. -27r122 ° --q08 ° -20n3. _-IC INB_ ENG EXPENDED _ROP -I 85. -qll8. -190. --418. S-II INSULATION PURGE 6A_ -Sq. -1 20. -54. -1 _0 ° S-II FROST -204. -_50. -204. -qSO. S-IVR FROST -91 . -2nO. -91. -200.

1ST FLT STAGF AT S-IC OE£OS 828697° 1826954. 8_7292. 1823R66- S-IC OTB_ ENGINE T-n- PROP -3668. -8fl87. -363q. -8011. S--ICIS-II ULLAGE RKT PROP -33. -73. -3T_. -73.

1ST FLT STAGE AT _IC/SII SFP 82q991° ]8197qq. 823625° 1815783. S-IC STAGE AT SEPARATION -16qRq7° -363q26. _-16q381. -3623qq. S-ICIS-II INTFRSTAGF SMRLL -61q° -1353. -61_. -1X53° _-ICPS-II ULLAGE RKT PR_P -83. -lBq. -8_. -184.

?ND FLT STAGE AT S--TI SSC GSSqq7. 1qSX8 _X2° 6585q7. lq518q7° S-II FUEL LEAn ]. 3° 3. 3. S-ICIS-II ULLAGE RKT PROP -188° -qlq. --18q. -qO6.

ZND FLT STAGE AT S-II TEN 659750. lqS"Sq 18. 658363. lqSlqql° S-IT T.R. PROPFLLANT -582. '- 128_. -587° -l?Bq° S-IT START TANK -11 . -_5. -11 . -25. S--Ic£S-II ULLAGE RKT PROP -313. -689, --309. -6827.

2NO FLT STAGF AT MATNSTAEE G58353° l'¢51q2O° 657q_9° 1qqgq50. S-II MAINSTAGE + VENTING -q37232. --gO 3q 32. -q_X. %q 9q ° -960110. LAUNCH ESCAPE SYSTEM -q051. -89 _0. -qOq2. -89| O° S-ICIS-II INT_RSTAGE LA_GF -3g69. --R750. -3982° -8779. S-II T.Do PROPELLANT -57° -176° -qg° -|09°

2No FLT STAGF AT S-IT C.O.S. ?130qq. qf;qGR2 . 21 ._8 BR ° q715q?. S-II T.D. PROPELLANT -160. -353° -128. -783 • S-IVn ULLAGE PROPELLANX -2. -5 ° -?. -5.

27ND FLT STG AT SII/SIV_ SFP 212R82. _Gq32q. 213757. q71?Sq. K-II STAGE AT SEPARATION --q27[1?° -9q] ql ° -q3q 36. -95759° S-III%-IVB INTERSTAGE-D_Y -3185 ° -TD 21 . -31 80. -70 10 o S-IIIS-IVR IS PROP -4 81 . i- 10 GO. -q 8q. -1066. S-IvB AFT FRAME -22 ° -'qB. -227. -q8. S-IvB ULLAGE PROPELLANT -1. -3. --1 ° -3. S-IvR DET PACKAGE -1 . -3. -I . -3.

2O-ll Table 20-9. Flight Sequence Mass Summary (Continued)

PRE D1 CTEf_ AC TU AL MASS HI%TO_Y K G LRM K G LBM

3RD FLT STG AT IST SSC lGG490. 3G7OqR. lGGG34. 367365. S-IVR ULLAGE PROPVLLAN? -qo. -R8. -qO. -88,, S-IvR FUEL LEAD LOSS -O. --O. -23. -50.

3RD FLT STG AT ]ST STVB IGN 1F, 6q 50. 366960. 16_;571. 367227, S-IVB ULLAGE PROPFLLANT - IO. - 22. - IN. -22. _-IvR START TANK -2. -4. --2. -_. S-IV_ T-B- PROPELLANT -lqB. -437. -IF,3. -3_0°

3RD FLT STG AT MAINSTAGE 1F,_240. 36G_q7. IGF,39G . 3GGBql° S-IVB ULLAGE ROCKET CASGS -G1. -! $5. -6! . -13q. S-IVq MAINSTAGE PROP -30424. -G7D73. -31152. -G8G78. %-IvB APS PROPELLANT -I . -_2. --I . -3.

3GO FLT STG AT ]ST SIVR COS 135755. 29928R. 135182° 298076. S-IVR I.D. PROPELLANT -i15. -49. -8I ° -]78°

3RD FLT GIG AT END ]ST ID 13570q° 2991RB. 135102. 2978_8. S-IVR ENG PqOP EXPENDFO -18. -40. -IR. --_O. S-IvR FUEL TANK LOSS -1 l 51 ° -25XR. --972. --2143. S-IVB LOX TANK LOSS -2D. '--4q. --4. --8. S--IvR APS PROPELLANT -I If2. -225. -51 . -I 13. S-TVR START TANK -1. -_2. -1. -2. S-IVR 021H2 BURNER -7 ° - IG . -7. -IS .

3RD FLT STG AT 2NO SSC 13q41D. 296323. 134048. 295._ 26 . S-IVB FUEL LEAD LOSS -2. -5° --2. -5.

3RD FLT STG AT 2NO SIVR IGN 13qq08° 296318. 1340_G. 295!;21 ° S-IVB START TANK -2 . -4 . -2 . -4 . %-IVR T.B. PROPELLANT -2 12. -q _,8 . -- 1Gl . -3SG.

3RD FLT STG AT MAINSTAGE 134194. 2'9_5846° 133883. 295151. S-IV_ MAINKTAGE PROP -7099q.. - l'r_G514. -706 8q. -155832,, S-IvR APS PROPELLANT -q. -8. --q. -20.

3RD FLT STG AT ?Nn STVB COS 63196. 139324° G31Bq. 1393(19o S-IVB T.n. PROPELLANT -45 . -I r'IO ° -74 . -I G3,,,

3RD FLT GIG AT END 2NO IO 63151° 17;9223. 63116.. 139146° JETTISON SLA -liDS. -2570. -liDS. -2571. COMMAND SERVICE MODULE -28639° --63679. --28806° --63507° S-IVB STAGF LOSS -53. -I 17. -559. -1232.

START OF TRANSIDOCKTNG 33093. 72457. 32584. 71836. COMMAND SERVICE MODULE 28839. '6 3579° 28806. 63507. S--IVR STAGE LOSS O. O. -O. -1.

END OF TRANSIDOCKING 61932° 136_36. 61390. 135342. COMMAND SERVICE MODULF -28839. -63579 . -2 8806 . -63507. LUNAR MODULE -ISll3. -33318. -1SD95. -33278° S--IVR STAGF LOSS O. O. -341 . -752.

LAUNCH VEH AT S/C SEPARATION 17980. 39GXq. 17148. 37R05° SPACECRAFT NOT SEPARATED --G2G. -1380. -62G. -1380. INSTRUMENT UNIT -I 953 ° - 4306 . -1939 ° --4275 . S-IVB STAGr AT SEPARATION -IS4Ol. -33953. -14583. -32150.

20-12 Table 20-10. Mass Characteristics Comparison

MASS LONGITUDINAL RADIAL ROLL MOMENT PITCH MOMENT YAW MOMENT C.G. (X STA.! C.G. OF INERTIA OF INERTIA OF INERTIA EVENT KILO 0/0 METERS METERS K G--M2 0/0 KG-M2 OlO KG--M2 0/0 POUNDS DEV. INCHES DELTA INCHES DELTA XlO-G DEV. X10-6 DEV. XlO-6 DEV.

1 30q75. 9.3G8 .f15 80 PRFD 288750. 3G8.8 2.2897 2.602 IG.6q8 1G.566 S-IC STAGE DRY 130422. 9.'368 .ODD °0580 • DO 00 AcTUAl[ 287531. -.92 368.8 .OO 2.2897 .OOO0 2.591 ".92 lfi.5 TR --.92 lG.qgfi -.q2

5262° 91.623 °T5qG PRFD 11600. ] 638°7 6.08 77 .139 .D 81 .081 S-ICIS-II INTER- --_- STAGE. TOTAL 5255. q 1°&26 °003 °1563 .0017 ACTUAE I1585. -.12 1638.8 .]O G.1555 .0678 .134 -.12 .080 -,.12 .081 -,,12

3625 I° 98. II 5 . 18 75 PRED 799I 8. 1894.3 7. 3829 .GOD 2.0 27 2.038 S-II STAGE. DRY I'D 36158. q8.DGq -.051 .1875 .OOO0 0 I ACTUAU 79719. -°25 1892.] -2.00 7°3829 °DO00 .597 -.50 1.997 -1°97 2,.,009 -l.qO

C_ 3GGG. 65-8G0 .0573 PRED 8081. 2592°9 2.2561 o065 .093 .Oq_ S-IIIS-IVB INTER- STAGE.TOTAL 3650. 65o936 .076 .0598 o_025 ACTUAU 8095° -.qq 25_5.9 3.00 2.3537 .0976 .065 -.94 .Oq3 -._q .Oqq -.q_

I13_0. 72.560 .?199 PRFO 25000. 2856.7 8.6377 .082 °298 .298 S-IVB STAGE.DRY 11273° 72°560 .OOO o2194 oDD00 ACTUAU 29852. -.59 2856.7 .DO 8._377 °ODD0 .081 -.59 .296 --.59 .296 -°59

1959. 82.915 .3576 PRED 306. 32eq .7 I q.fiR 1!1 .0 19 .'0 lO . 009 VEHICLE INSTRUMENT- UNIT lqqO. 82°_15 .ODD .3570 -.0007 ACTUAL _275. -.71 3299°7 °O01q. I1595 -. 0256 .019 -.71 .lOlO -.71 °009 -.71

q8731. 91.653 .1085 PREO 107q33. 3GOS.W q.2720 .ngO 1.552 1.555 SPACECRAFT, TOTAL 9862&. 91o&58 .DOS .1099 °_Olq ACTUAU 107200. -.21 3608.6 .20 q°3267 .05q7 .088 -1.70 1.5q9 -.21 l._O -.30 L8 ° 9_6° _ L8 ° Z£6°ht £_'- [L8" OOIO'- _OSI'Z L£'£ - 8"5bLZ Otl" "StIS IL_ NVN13V £001)'- 9hSO" 580 °- _IU'IL "068£IZ 3VN9IS 3JOLR3 IV 39¥15 IH9I]_ ON_ 0_5. t_b gt;S=tb SLS" Zn 91 °Z Z "66LZ "Z8969_ 03Od 6tl SO" 66D" [L

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(panu.L:_uo3) UOS.LJ_dLu03SD.L_S._Ja_3_J_q3 ss_H "Ot-O_ aLqEl Table 20-10. Mass Characteristics Comparison (Continued)

MASS LONGITUDINAL RADIAL ROLL MOMENT PITCH NOMENT TAW MOMENT CoG. (X STA.! C.G. OF INERTIA OF INERTIA OF INERTIA EVENT ------KILO 0/0 METERS METERS KG--M2 010 KG-M2 OlD KG-M2 -010 POUNDS DEV. INCHES DELTA INCHES DELTA X10-6 OEV. XI0-6 DEV. X10--8 DEV.

I 34908. 78.053 ._4GG PRED 29G319. 3073.0 1.8350 .195 I2.608 I2.G08 3R0 FLIGHT STAGE AT 2ND IGNITION 1_047. 78.057 .004 .r1_74 °0007 ACTUAU 295521. -.2G 3073.1 .IG Io8E43 .029X .19G .41 12.591 -.13 12.590 -.14

13e194. 78.060 .0466 PRFD 295846. 3073.2 1.8350 .195 12.'G03 12.603 3RD FLIGHT STAGE AT 2ND MAINSTAGE 133883. 78.062 .002 o0479 .00f17 ACTUAb 2951G1. -.23 3073.3 .08 1.8E43 .029_ .19G .41 12.587 -.17 12.587 -.13

63 197. 85.770 .0975 3RD FLIGHT STAGE PRFD 1 39324. 3376.8 3.8399 .195 5 .,272 5.272 AT 2ND CUTOFF 0 SIGNAL 63190. 85.712 -.058 .119 89 .I10 14 I ACTUAU 139309. -.01 3374.5 -2.2_1 3.8948 .0559 .195 .¢15 5.329 I.08 5.327 l.OG

6315 1. 85.78] .0975 3RD FLIGHT STAGE PRFO I 39223. 3377.2 3.8394 .195 5,,261 5.260 AT 2ND END THRUST OECAY 63116. 85.731 -.OSO .0989 .ODlq ACTUAU 13914G. --.05 3375.2 -1.98 3.8998 .0554 .195 .95 5.309 .93 5.308 .91

33093. 78.781 .0825 PRFD 72957. 3 I0 I.G 3.2468 • 1 39 1.687 1=G84 CSM SEPARATED 32585. 78.799 .017 .0792 -',,,0033 ACTUAL 7183G. -1.53 3102.3 .G8 3. I1Gq -.13C4 .139 1.68G -.01 1.G82 -.10

61932. 85.218 . I292 PRFD 1 36536. 3355.1 5.0850 4.:719 4.715 Cq;M DOCKEn 61391. 85.267 .099 .I277 --. O0 14 ACTUAU 135342. -,,.87 335,7.0 1.92 5.0293 -.Q557 .18G .30 4.691 -o59 4.G8£ -.El

17980. 73.615 .1517 PRED 39639. 2898.2 5.9718 .109 .,G 14 .G 11 SPACECRAFT SEP- ARATED 17149° 73.573 -.042 .145G -.00GI ACTUAL 37805. -4.62 2896.6 -l.G6 5.7328 -.2390 .109 .06 .GO9 -°8| .605 -1.08 SECTION 21

MISSION OBJECTIVES ACCOMPLISHMENT

Table 21-I presents the MSFC Principal Detailed Objectives and Secondary Detailed Objectives as defined in the Saturn V Mission Implementation Plan, Mission G, Revision C. An assessment of the degree of accomplishment of each objective is shown. Discussion supporting the assessment can be found in the indicated sections of the Saturn V Launch Vehicle Flight Evaluation Report - AS-506, Apollo II Mission.

Table 21-I. Mission Objectives Accomplishment Summary

MSFC PRINCIPAL DETAILED DEGREE PARAGRAPH NO. OBJECTIVES (PDO) AND SECONDARY OF DISCREPANCIES IN WHICH DETAILED OBJECTIVES (SDO) ACCOMPLISHMENT DISCUSSED

4.1 Launch on variable 72 to 108-degree Complete None 4.3.1 flight azimuth and insertion of 4.3.2 S-IVB/IU/SC into a circular earth 11.4.2 parking orbit (PDO). 4.1 4.3.3 Restart the S-IVB during either the Complete None 7.6 second or third revolution and in- 10.3 jection of the S-IVB/IU/SC onto the II .4.4 planned translunar trajectory (PDO).

Provide the required attitude control Complete None 4.1 for the S-IVB/IU/SC during the Trans- 4.3.4 10.3 position, Docking, and Ejection (TD&E) maneuver (PDO). II .4.4

Use residual S-IVB propellants and Complete None 4.3.5 Auxiliary Propulsion System (APS), 7.13 after final LV/SC separation, to 10.3 safe the S-IVB and to minimize the II .4.4 possibility of the following, in order of priority: I. S-IVB/IU recontact with SC 2. S-IVB/IU earth impact 3. S-IVB/IU lunar impact (SDO).

21 -I/21-2

SECTION22

FAILURES,ANOMALIESANDDEVIATIONS

22.1 SUMMARY

Evaluation of the launch vehicle performance during the AS-506 flight revealed no failures or anomalies and ten deviations. None of these deviations had an adverse effect on the mission.

22.2 SYSTEMFAILURESANDANOMALIES

There were no failures or anomalies detected during the launch vehicle operational period of flight. 22.3 SYSTEMDEVIATIONS

Ten system deviations occurred, none of which had any significant effect on the flight or operation of the particular systems involved. Table 22-I presents these deviations along with the corrective actions being considered and references to paragraphs containing additional discussion of the deviations.

22-I Table 22-I. Summary of Deviations

CORRECTIVE ACTION PARAGRAPH VEHICLE DEVIATION PROBABLE CAUSE BEING CONSIDERED REFERENCE SYSTEM 5.6 S-IC Unexplained LOX suction Unknown. None. Similar occurrences Propulsion duct pressure decay of during AS-503, AS-504 and engine No. 5 after Center AS-505 with no effect on Engine Cutoff (CECO). mission.

3.6.2 S-II J-2 engine No. 1 start Lower than planned Ground Increase GSE regulator nomi- Propulsion tank pressure below pre- Support Equipment (GSE) nal setting and relax pre- 6.2 launch commit (-33 regulator setting. launch commit redline to seconds) redline, more closely approximate actual requirements.

6.2 S-II J-2 engine No. 2 helium Leakage through engine helium None. Decay rate returned Propulsion tank pressure decay rate regulator. to normal at 30 seconds sharper than expected after ESC. after Engine Start Command (ESC).

8.4 S-IVB I. S-IVB engine driven I. Inherent "drift-up" of Under investigation. Hydraulics hydraulic pump system pump plus uncompensated pressure drifted thermal expansion in com- 16 N/cm2 (23.2 psi) pensator; neither of which over the predicted was included in establish-i upper limit of ing the predicted upper 2526 N/cm 2 (3665 psia) limit. at 9848 seconds. 2. Later exhibited small 2. Abrupt change could be due but abrupt drop in to frictional hysteresis pressure. in the pressure/flow regulating mechanism.

7.10.2 S-IVB LOX tank pressure decayed I. Thermal collapse of None. Probably due to thermal collapse. Since LOX Propulsion approximately 5 percent ullage pressure. below predicted minimum tank repressurization was during coast in earth 2. LOX tank leakage. well within design capa- parking orbit. bilities, this pressure decay was not considered a problem (even if second opportunity restart had been required).

None. While this is appar- 9.2.3.1 S-IVB Low amplitude, 17 to 20 Similar to oscillations on Structures hertz longitudinal oscilla- AS-505, but only 20 percent rently a phenomenon which tions during first burn. of the amplitude. Data in- is characteristic of the dicates typical buildup and stage, changes in the engine decay periods of very mild or payload configuration oscillations without indi- would require a reassessment. cations of propulsion/structural coupling.

Improve scaling in IU LVDC 6.5 Instrument Delay of 6 seconds in IU Primarily due to improper Unit (IU) command to shift S-II Engine scaling in IU LVDC velocity velocity calculations. 10.2.1 Guidance Mixture Ratio (EMR). computations. 10.3

Raise the maximum pressure 18,4.2 IU/Gas Bearing Inertial platform gas bear- Inherent in the system late differential spec. to an Supply ing differential _ressure in the flight. Occurred on drifted 0.54 N/cm z (0.8 psi) AS-502, AS-503, AS-504 and acceptable value. Ground above specification at AS-506 and caused no problem. tests indicate no performance deviations at 23,200 seconds. 13.8 N/cm 2 (20 psi).

Under investigation, but had 10.2 IU/ST-124 ST-124 platfom crossrange Vibration caused the Y accel- no effect on operation of Inertial velocity exhibited negative erometer to have a level launch vehicle. I0.4.7 Platform 1.8 m/s (5.9 ft/s) shift shift or to touch a mechani- 3.3 seconds after liftoff. cal stop.

Malfunction of coaxial switch. Coaxial switch has been re- 19.4.3.2 IU/RF Erratic signal strength at receiving station be- placed on AS-507 with new ginning at 27,128 seconds. design.

22 -2 SECTION 23

SPACECRAFT SUMMARY

The purpose of the Apollo II mission was to land men on the lunar surface and to return them safely. The crew was Neil A. Armstrong, Commander; Michael Collins, Command Module (CM) Pilot; and Edwin E. Aldrin, Jr., Lunar Module (LM) Pilot.

The space vehicle was launched from Kennedy Space Center (KSC), Florida, at 9:32:00 a.m. Eastern Daylight Time (EDT), July 16, 1969. The activities during earth orbit checkout, Translunar Injection (TLI), transposition and docking, spacecraft ejection, and translunar coast were similar to those of Apollo I0. Only one midcourse correction, performed at about 27 hours Ground Elapsed Time (GET), was required during translunar coast.

The spacecraft was inserted into lunar orbit at approximately 76 hours, and the circularization maneuver was performed two revolutions later. Initial checkout of LM systems was satisfactory, and after a planned rest period, the Commander and LM Pilot entered the LM to prepare for descent.

The two spacecraft were undocked at I00 hours, followed by separation of the Command and Service Modules (CSM) from the LM. Descent orbit insertion was performed at about 101.5 hours, and powered descent to the lunar surface began about 1 hour later. Operation of the guidance and descent propulsion systems was nominal. During the final 2.5 minutes of descent, the LM was maneuvered manually approximately 305 meters (I000 ft) downrange. The spacecraft landed in the Sea of Tranquility at 102:45:40. The landing coordinates were 0.647 degree north latitude and 23.505 degrees east longitude, based on identification of landmarks from the onboard sequence camera. During the first 2 hours on the surface, the two crewmen performed a postlanding checkout of all LM systems. Afterwards they ate their first meal on the moon and elected to perform the surface operations earlier than planned.

Considerable time was devoted tocheckout and donning of the back-mounted portable life support and oxygen purge systems. The Commander egressed through the forward hatch and deployed an equipment module in the descent stage. A camera in this module provided live television coverage of the Commander descending the ladder to the surface, with first contact made

23-I at 109:24:19 (10:56:19 p.m. EDT, July 20, 1969). The LM Pilot egressed soon thereafter, and both crewmen used the initial period on the surface to become acclimated to the reduced gravity and new surface conditions. A contingency sample was taken from the surface, and the television camera was deployed so that most of the LM was included in its view field. The crew took numerous photographs, erected a U.S. flag, and activated the scientific experiments, which included a solar wind detector, a passive seismometer, and a laser reflector. The LM Pilot spent considerable time evaluating his ability to operate and move about, and despite the limitations imposed by the pressurized suit, he was able to translate rapidly and with confidence. Approximately 24 kilograms (54 Ibm) of bulk surface material were collected to be returned for analysis. The crew reentered the LM at 111:39:00, with surface exploration lasting 2 hours, 31 minutes.

Ascent preparation was conducted efficiently, and the ascent stage lifted off the surface at 124.5 hours A nominal firing of the ascent engine placed the vehicle into an 83 by 17 kilometer (45 by 9 n mi) orbit. After a rendezvous sequence similar to that of Apollo I0, the two spacecraft were docked at 128 hours. Following transfer of the crew, the ascent stage was jettisoned, and the Commandand Service Modules were prepared for transearth injection.

The return flight started with a 150-second firing of the service propulsion engine during the 31st lunar revolution at 135.5 hours. As in translunar flight, only one midcourse correction was required, and passive thermal control was exercised for most of transearth coast. The possibility of inclement weather necessitated moving the landing point 398 kilometers (215 n mi) downrange. The entry phase was normal, and the CM landed in the Pacific Ocean at 195:18:35. The landing coordinates, as determined from the onboard computer, were 13.3 degrees north latitude and 169.4 degrees west longitude. After landing, the crew donned biological isolation garments and were re- trieved by helicopter and taken to the primary recovery ship, USSHornet. The crew then entered the Mobile Quarantine Facility, which arrived at the Lunar Receiving Laboratory in Houston on Sunday, July 27, 1969. The CMwas taken aboard the Hornet about 3 hours after landing. The lunar samples arrived at the Receiving Laboratory the day after landing.

For further details on the spacecraft performance, refer to the Apollo II Mission Report published by NASAManned Spacecraft Center at Houston, Texas.

23-2 APPENDIXA

ATMOSPHERE

A.I SUMMARY

This appendix presents a summary of the atmospheric environment at launch time of the AS-506. The format of these data is similar to that presented on previous launches of •Saturn vehicles to permit comparisons. Surface and upper winds, and thermodynamic data near the launch time are given.

A.2 GENERAL ATMOSPHERIC CONDITIONS AT LAUNCH TIME

A high pressure cell, in the Atlantic Ocean off the North Carolina coast, along with a weak trough of low pressure located in the northeastern Gulf of Mexico caused light southerly surface winds and brought moisture into the Cape Kennedy, Florida area, which contributed to the cloudy conditions and distant thunderstorms that were observed during launch.

A.3 SURFACE OBSERVATIONS AT LAUNCH TIME i At launch time, total sky cover was 9/10 with I/I0 cumulus at 0.7 kilometer (2400 ft), 2/10 altocumulus at 4.6 kilometers (15,000 ft) and 9/10 cirrostratus at an unknown altitude. Surface observations at launch time are summarized in Table A-I. Solar radiation data are given in Table A-2.

A.4 UPPER AIR MEASUREMENTS

Data were used from three of the upper air wind systems to compile the final meteorological tape. Table A-3 summarizes the data systems used.

A.4.1 Wind Speed

Wind speed was light in the lower levels. In the maximum dynamic pressure region a peak speed of 9.6 m/s (18.7 knots) was observed at 11.40 kilometers (37,400 ft). At higher altitudes the wind speed increased steadily, as shown in Figure A-I.

A.4.2 Wind Direction

The surface wind was from the south, but with altitude shifted clockwise through west, north and then stayed easterly above 16 kilometers (52,490 ft) altitude, as shown in Figure A-2.

A-I Table A-I. Surface Observations at AS-506 Launch Time

WIND TIME PRES- TEM- DEW VlSI- AFTER SURE_ PERATURE POINT BILITY AMOUNT SKY COVER HEIGHT SPEED DIR LOCATION T-O N/CMz °K °K KM (TENTHS) TYPE OF BASE M/S (DEG) (MIN) (PSlA) (°F) (°F) (STAT MI) (KNOTS)

i 180 Kennedy Space 0 10.203 302.6 297.0 16 1 Cumulus 700 1.0 Center, Station (14.80) (85.0) _75.0) (I0) (2400) (2.0) Merritt Island, 2 iAlto- 4600 :lorida cumulus (15,000) 9 Cirro- high stratus

18O Cape Kennedy 13 10,195 303.0 297.5 ...... 1.0 Rawinsonde (14.79) (85.6) 75.7) (2.0) Measurements 175 Pad 39A Lightpole 0 ...... 3,3 SE 18.3 m * (6.4) (60.0 ft)

*Above Natural Grade

A.4.3 Pitch Wind Component

The surface pitch wind speed component was a tail wind of 0.3 m/s (0.6 knots). A maximum tail wind of 7.6 m/s (14.8 knots) was observed at 11.18 kilometers (36,680 ft) altitude. Head winds were observed above 15.0 kilometers (49,210 ft) altitude. See Figure A-3o

A.4.4 Yaw Wind Component

The yaw wind speed component was a wind from the right at the surface to approximately 9.0 kilometers (29,530 ft) altitude. Winds from the left prevailed above this altitude to 16.3 kilometers (53,480 ft) with a peak yaw wind speed of 7.1 m/s (13.8 knots) at 12.1 kilometers (39,530 ft) altitude. Above 16.3 kilometers (53,480 ft) yaw winds were from the right. See Figure A-4.

A.4.5 Component Wind Shears

The largest component wind shear (Ah = I000 m) in the altitude range of 8 to 16 kilometers (26,247 to 52,493 ft) was a pitch shear of 0.0077 sec -I at 14.8 kilometers (48,490 ft). The largest yaw wind shear, in the lower levels, was 0.0056 sec-I at 10.3 kilometers (33,790 ft). See Figure A-5.

A-2 Table A-2. Solar Radiation at AS-506 Launch Time, Launch Pad 39A

TOTAL NORMAL DIFFUSE HOURENDING HORIZONTAL INCIDENT SKY DATE EST G-CAL/CM2 G-CAL/CM2 G-CAL/CM 2 (MIN) (MIN) (MIN)

July 15, 1969 0600 0o00 0.00 0.00 0700 0.II 0.03 0.I0 O8OO 0.15 0.02 0.14 0900 0.21 0.02 0.20 I000 0.41 0.03 0.39 II00 0.57 0.03 0.54 1200 0.78 0.12 0.66 1300 1.17 O.44 0.74 1400 0.89 0.29 0.62 1500 0.39 0.02 0.37 1600 0.33 0.01 0.32 1700 0.43 0.06 O. 40 1800 0.30 0.I0 0.27 1900 0.07 0.00 0.07 2000 0.01 0.00 0.01

july 16, 1969 060O 0.01 0.00 0.01 0700 0.14 0.II 0.II 0800 0.42 0.36 0.24 0900 0.77 0.54 O. 40 I000 0.88 0.37 0.57 II00 1.46 0.52 0.98

A.4o6 Extreme Wind Data in the High Dynamic Region

A summa_ of the maximum wind speeds and wind components is given in Table A-4. A summary of the extreme wind shear values is given in Table A-5.

A.5 THERMODYNAMIC DATA

Comparisons of the thermodynamic data taken at AS-506 launch time with the Patrick Reference Atmosphere, 1963 (PRA-63) for temperature, density, pressure, and Optical Index of Refraction are shown in Figures A-6 and A-7 and discussed in the following paragraphs.

A.5.1 Temperature

Atmospheric temperature deviations were small, being less than 3 percent deviation from the PRA-63. At most altitudes, the temperature was warmer than the PRA-63.

A-3 SON033$ '3WI1 3ONV'd

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RANGETIME, SECONDS "rl

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2 _. 1 30 20 0 0 .004 .008 .012 .016 0 .004 .008 .012 .016

(Sx) PITCH WIND SHEAR ( _F'!"sec-l) (S z) YAW WIND SHEAR(_.Aw sec_l)

Figure A-5. Pitch (Sx) and Yaw (Sz) Component Wind Shears At Launch Time of AS-506

A-8 Table A-3. Systems Used to Measure Upper Air Wind Data for AS-506

RELEASE TIME PORTION OF DATA USED

START END TIME TYPE OF DATA TIME AFTER TIME TIME (UT) T-O ALTITUDE ALTITUDE AFTER AFTER (MIN) M M T-O T-O (FT) (FT) (MIN) (MIN)

I i

FPS-I 6 Jimsphere 1347 15 0 15 16,250 7O (53,310)

Rawinsonde 1345 13 16,500 67 24,750 94 (54,130) (81,200)

Loki Dart 1512 I00 56,000 I01 25,000 124 (183,725) (82,020)

Table A-4. Maximum Wind Speed in High Dynamic Pressure Region for Apollo/Saturn 501 through Apollo/Saturn 506 Vehicles

MAXIMUM WIND MAXIMUM WIND COMPONENTS VEHICLE SPEED ALT ALT YAW (Wz) ALT NUMBER DIR PITCH (Wx) M/S KM M/S KM M/S KM (BEG) (KNOTS) (FT) (KNOTS) (KNOTS) (FT)

i I AS-501 26.0 273 11.50 24.3 11.50 12.9 9.00 (50.5) (37,700) (47.2) (37,700) (25.1) (29,500)

AS-502 27.1 255 12.00 27.1 12.00 12.9 15.75 (52.7) (42,600) (52.7) (42,600) (25.1) (51,700)

AS-503 34.8 284 15.22 31.2 15.10 22.6 15.80 (67.6) (49,900) (60.6) (49,500) (43.9) (51,800) 11.70 21.7 11.43 IAS-504 76.2 264 11.73 74.5 (148.1) (38,480) (144.8) (38,390) (42.2) (37,500)

AS-505 42.5 270 14.18 40.8 13.80 18.7 14.85 (82.6) (46,520) (79.3) (45,280) (36.3) (48,720)

AS-506 9.6 297 II .40 7.6 11.18 7.1 12.05 (18.7) (37,400) (14.8) (36,680) (13.8) (39,530)

A-9 i

Table A-5. Extreme Wind Shear Values in the High Dynamic Pressure Region for Apollo/Saturn 501 through Apollo/Saturn 506 Vehicles

(Ah = I000 m)

PITCH PLANE YAW PLANE

VEHICLE NUMBER ALTITUDE ALTITUDE SHEAR SHEAR KM KM (SEC-I) (SEC-I) (FT) (FT)

AS-501 0.0066 I0.00 0.0067 I0.00 (32,800) (32,800)

AS-502 0.0125 14.90 0.0084 13.28 (48,900) (43,500)

AS-503 0.0103 16.00 0.0157 15.78 (52,5OO) (51,800)

AS-504 0.0248 15.15 0.0254 14.68 (49,700) (48,160)

AS-505 0.0203 15.30 0.0125 15.53 (5O,2OO) ( 5O ,950)

AS-506 0.0077 14.78 0.0056 10.30 (48,490) (33,790)

A.5.2 Atmospheric Pressure

Atmospheric pressure deviations remained greater than the PRA-63 values at all altitudes. Surface pressure was 0.2 percent greater than the PRA- 63 and increased to a peak deviation of 9.0 percent at 44.0 kilometers (144,360 ft).

A.5.3 Atmospheric Density

Atmospheric density deviations were small, being less than 5 percent deviation from the PRA-63 from the,surface to 29.8 kilometers (97,770 ft) altitude. Density deviations increased above this altitude and reached a peak of 10.3 percent at 46.0 kilometers (150,920 ft). Surface atmospheric density was -2.1 percent of the PRA-63 surface density.

A-IO 6O

54 / (

5O f 48 ( 46 / 44 \ / 42 k 40 f i 38 > 36 / 34 o

__3o / I-- _ 28

26 f

24 <

22 < ( 20 J 18

16 ( 14 P

8 / L_ 6 4 < 2 ! I 0 -4 -2 0 2 4 6 -6 -4 -2 0 2 4 6 8 I0 12

RELATIVE DEVIATION OF TEMPERATURE,PERCENT RELATIVE DEVIATION OF DENSITY, PERCENT

Figure A-6. Relative Deviation of Temperature and Density From the PRA-63 Reference Atmosphere, AS-506

A-II 60

56 150

48 140

130 40

34 12o

32 8

110 _ 26

22 100

90 16

8O 12

4 5O

4O 2 ' 30 20 0 -4 -2 0 2 4 6 8 I0 12 14 -14 -12 -I0 -8 -6 -4 -2 0 2

RELATIVE DEVIATION OF PRESSURE, PERCENT ABSOLUTE DEVIATION OF THE OPTICAL INDEX OF REFRACTION 10 -6

Figure A-7. Relative Deviation of Pressure and Absolute Deviation of the Index of Refraction From the PRA-63 Reference Atmosphere, AS-506

A-12 A.5.4 Optical Index of Refraction At the surface, the Optical Index of Refraction was 12.9 x 10-6 units lower than the corresponding value of the PRA-63. The deviation became less negative with altitude, becoming a maximumpositive deviation of 2.43 x 10-6 greater than the corresponding value of the PRA-63 at 14.3 kilometers (46,920 ft). Above this altitude the Optical Index of Re- fraction approximates the PRA-63 values. A.6 COMPARISONOF SELECTEDATMOSPHERICDATAFORSATURNV LAUNCHES

A summary of the atmospheric data for each Saturn V launch is shown in Table A-6.

Table A-6. Selected Atmospheric Observations for Apollo/Saturn 501 Through Apollo/Saturn 506 Vehicle Launches at Kennedy Space Center, Florida

VEHICLE DATA SURFACE DATA INFLIGHT CuNDITIONS

VEHICLE DATE TIFIE LAUNCH PRESSURE TEMPERA- RELATIVE WI_4D* CLOUDS MAXIMUM WIND IN 8-16 KM LAYER NUMBER NEAREST COMPLEX N/CM 2 TURE °C HUMIDITY SPEED DIRECTION ALTITUDE SPEED DIRECTION MINUTE PERCENT M/S DEG KM M/S DEG

AS-501 9 Nov 67 0700 EST 39A 10.261 17.6 55 8.0 70 I/IO cumulus 11.50 26.0 273

20.9 83 5.4 132 5/10 stratocumulus ]3.00 27.1 255 AS-502 4 Apr 68 0600 EST 39A 10.200

AS-503 21 Dec 68 0751 EST 39A 10.207 15.0 88 1.0 360 4/10 cirrus 15.22 34.8 284

AS-504 3 Mar 69 II00 EST 39A 10.095 19.6 61 6.9 160 lO/10 strato- 11.73 76.2 264 cumulus

26.7 75 8.2 125 4/10 cumulus, 2/70 14.18 42.5 270 AS-5D5 18 May 69 1149 EDT 39B 10.190 a3tocumulus, 10/10 cirrus

11.40 9.6 297 AS-506 16 Ju} 69 0932 EDT 39A 10.203 29.4 73 3.3 175 I/lO cumulus, 2/10 altocumulus, 9/10 cirrostratus

*Instantaneous readings from charts at T-O from anemometers on launch pad at 18.3 m (60.0 ft) on launch complex 39 (A&B). Heights of anemometers are above natural grade.

A-13/A-14

APPENDIXB

AS-506 SIGNIFICANT CONFIGURATIONCHANGES

B.I INTRODUCTION

AS-506, sixth flight of the Saturn V series, was the fourth manned Apollo Saturn V vehicle. The AS-506 launch vehicle was configured the same as the AS-505 with significant exceptions as shown in Tables B-I through B-4. The basic AS-506 Apollo II spacecraft structure and components were unchanged from the AS-504 Apollo 9 configuration except lunar module crew provisions were accompanied by portable life support systems and associated controls required to accommodate extra vehicular surface activity. The basic vehicle description is presented in Appendix B of the Saturn V Launch Vehicle Flight Evaluation Report AS-504, Apollo 9 Mission, MPR-SAT-FE-69-4.

B-I Table B-I. S-IC Significant Configuration Changes

SYSTEM CHANGE REASON

Control Deleted prevalve accumu- Stage system tests have shown Pressure lator bottles. that the accumulator bottles are not required for satisfactory closure of prevalves.

Data Measurements reduced from R&D instrumentation which is 669 to 313; deletions no longer required. include all vibration and acoustic measurements. Deleted 3 PAM, 2 FM/FM, Deletion of R&D instrumenta- and 2 SS/FM systems. tion permitted reduction of telemetry system. Deleted airborne tape R&D data recording system recorder. which is no longer required. Modified register'switches Improve reliability. card in telemetry PCM/DDAS assembly.

Electrical Capacity of instrumenta- Deletion of R&D instrumentation battery ID20 reduced. tion permitted use of lower capacity battery.

Table B-2. S-II Significant Configuration Changes

SYSTEM CHANGE REASON

Instrumentation Reduction of measurement Maturity of designoR&D quantity from 1018 on instrumentation no longer S-II-5 to 563 on S-II-6 required. and subs. Deleted 3 PAM, l FM/FM, and 2 SS/FM systems. Deleted 2 airborne tape recorders. Thermal Deleted electronic pack- Instrumentation reduction. Control ages 206A84, 206A85, 208, 211, 212, 213 from aft skirt area and pack- ages 222, 224, 227, and 228 from forward skirt area.

Structures Deleted Ill-inch dollar Analysis and tests indicated weld doublers on aft the doublers not necessary. LOX bulkhead.

B-2 Table B-3. S-IVB Significant Configuration Changes

SYSTEM CHANGE REASON

Instrumentation Five S-IVB measurements To better define the low are routed through the frequency vibration which IU/FM/FM telemetry system. occurred on AS-505. (Remaining measurements same as AS-504).

Propulsion Addition of liner to LH2 To eliminate flow resonance feed duct. problems. 02/H2 injector change. To eliminate possible burn through during flight operation. Addition of block point To prevent possible over- to shutoff valve of heating of solenoid in pneumatic power control secondary regulation mode module. (bang-bang).

New configuration cold To prevent main poppet seat helium shutoff valves distortion at low tempera- for cryogenic repress ture. appl i cati on. New configuration cold To prevent main poppet seat helium dump valve. distortion at low temperature.

Thermal protection- High solenoid cold tempera- pneumatic shutoff valve ture in bang-bang mode of solenoid. operation will reduce the solder strength.

B-3 Table B-4. IU Significant Configuration Changes

SYSTEM CHANGE REASON

Environmental "Tee" section added to ends Provide capability for RTG Control of air/GN2 purge duct. fuel cask preflight thermal "Tee" is capped on AS-506. conditioning. Additional ducting, nozzle and brackets used on AS-505 not included on AS-506.

Thermal switch settings Settings determined from test were: data. Open at 288.8°K(60.I°F) Close at 288.2°K(59.0°F)

Additional clamp added to AS-505 preflight thermal IU air/GN2 purge duct boot conditioning to RTG was lost at the umbilical plate. during countdown. Suspect area Increased torque on clamps was the clamp at the inlet to associated with the duct the IU. boot.

Instrumentation Two acceleration and three Low frequency structural and pressure measurements added vibrations monitored during Communication to the S-IVB are telemetered the AS-506 flight. via IU FM/FM system. Added measurements:

A12-403 Gimbal block longitudinal accelerometer A15-424 LOX feedline at Aft LOX dome accelerometer. DI-401 Thrust chamber pressure. D3-403 Oxidizer pump inlet pressure. D9-401 Oxidizer pump discharge pressure. Networks Additional Cables and modifica- Modifications required to add five tions to the measuring distrib- measurements for the S-IVB. utor and F1TM assembly.

Flight Launch pad choice from target Pad choice can be loaded with target- Program tape. ing parameters from tape via the RCA- IIOA. Eliminates necessity for re- assembly of flight program due to change of launch pad.

Capability for detection of An S-IC engine out, formerly not early S-IC engine out. detectable until 14 seconds after liftoff, can now be detected from 6 seconds after liftoff.

Accelerometer zero test. Adjusts the Sin D term for an early S-IC or S-II outboard engine out.

Expanded S-II IGM guidance. Automatically adjusts for wide variations in performance, for either high or low thrust levels. Has inherent capability to adjust guidance for multiple S-II engines out.

Deletion of program recogni- The program will not prevent a time base tion of critical pairs of update from altering the time separation switch selector commands. between any pair of switch selector commands. It will be the responsibility of ground controllers to maintain such requirements if they exist.

B-4 Table B-4. IU Significant Configuration Changes (Continued)

SYSTEM CHANGE REASON

Flight Expanded data compression Maximum duration of data compression Program (Cont'd) capability. period extended from 50 to 95 minutes. Sample rate of Table 3 changed from 30 to 60 seconds. Maximum sample capacity of Table 2 increased from 31 to 59; Table 4 increased from 61 to 116.

Deletion of LH2 propel- The S-IVB residual LH2 dump was deleted, lant dump. since velocity change requirements could be satisfied otherwise. Selective telemetry cali- Capability is provided to distinguish bration and dump. between "dump and calibrate" and "calibrate only" LVDC telemetry stations. Of the 14 LVDC telemetry stations on the mission, onlyCarnarvon, Hawaii, and Guaymas are compressed data dump stations.

MSFC--RSA, Ala B-5/B-6

MPR-SAT-FE-69-9

APPROVAL

SATURN V LAUNCH VEHICLE FLIGHT EVALUATION REPORT

AS-506, APOLLO II MISSION

By Saturn Flight Evaluation Working Group

The information in this report has been reviewed for security classification. Review of any information concerning Department of Defense or Atomic Energy Commission programs has been made by the MSFC Security Classification Officer. The highest classification has been determined to be unclassified.

Stanley L. SFragge - I Security Classification Officer

This_report has been reviewed and approved for technical accuracy.

George'. McKay, Jr./ ' / Chairman, Saturn Fl_ht EValuation Working Group

/_ Ire_c_ori_Sclence" " and En-1g'neering

Roy E. Godi_ey Saturn Program Manager DISTRIBUTION:

MSFC: S&E-AERO Mr. Hellebrand, S&E-ASTN-DIR Mr. Edwards, S&E-ASTN-DIR Dr. von Braun, DIR Dr. Geissler, S&E-AERO-DIR Mr. Sterett, S&E-ASTN-A Mr. Shepherd, DIR Mr. Horn, S&E-AERO-DIR Mr. Schwinghamer, S&E-ASTN-M Dr. Rees, DEP-T Mr. Dahm, S&E-AERO-A (2) Mr. Earle, S&E-ASTN-P Mr. Gorman, DEP-M Mr. Holderer, S&E-AERO-A Mr. Reilmann, S&E-ASTN-P Dr. Stuhlinger, ADIR-S Mr. Dunn, S&E-AERO-ADV Mr. Thompson, S&E-ASTN-E Mr. Elkin, S&E-AERO-AT Mr. Fuhrmann, S&E-ASTN-EM E Mr. Wilson, S&E-AERO-AT Mr. Cobb, S&E-ASTN-PP (2) Mr. Jones, S&E-AERO-AT Mr. Black, S&E-ASTN-PPE Mr. Maus, E-DIR Mr. Reed, S&E-AERO-AU Mr. Wood, S&E-ASTN-P Mr. Smith, E-S Mr. Guest, S&E-AERO-AU Mr. Hunt, S&E-ASTN-A Mr. Ryan, S&E-AERO-DD Mr. Beam, S&E-ASTN-AD PA Mr. Cremin, S&E-AERO-M Mr. Riquelmy, S&E-ASTN-SDF Mr. Lindberg, S&E-AERO-M (I0) Mr. Katz, S&E-ASTN-SER Mr. Slattery, PA-DIR Mr. Baker, S&E-AERO-G Mr. Showers, S&E-ASTN-SL Mr. Jackson, S&E-AERO-P Mr. Frederick, S&E-ASTN-SS PD Mr. Cummings, S&E-AERO-T Mr. Furman, S&E-ASTN-AA Mr. O. E. Smith, S&E-AERO-Y Mr. Green, S&E-ASTN-SVM Dr. Lucas, PD-DIR Mr. J. Sims, S&E-AERO-P Mr. Grafton, S&E-ASTN-T Mr. Williams, PD-DIR (2) Dr. Lovingood, S&E-AERO-D Mr. Marmann, S&E-ASTN-VAW Mr. Driscoll, PD-DIR Mr. Vaughan, S&E-AERO-Y Mr. Lutonsky, S&E-ASTN-VAW Mr. Thomason, PD-DO-DIR Mr. Devenish, S&E-ASTN-VNP (2) Mr. Goerner, PD-DO S&E-CSE Mr. Sells, S&E-ASTN-VO0 Mr. Nicaise, PD-DO Mr. Schulze, S&E-ASTN-V (2) Mr. Jean, PD-RV Dr. Haeussermann, S&E-CSE-DIR Mr. Rothe, S&E-ASTN-XA Mr. Digesu, PD-DO-E Mr. Hoberg, S&E-CSE-DIR Mr. Griner, S&E-ASTN-XSJ Mr. Palaoro, PD-SS Mr. Mack, S&E-CSE-DIR Mr. Boone, S&E-ASTN-XEK Mr. Blumrich, PD-DO-SL Dr. McDonough, S&E-CSE-A Mr. Aberg, S&E-CSE-S S&E-QUAL Mr. Fichtner, S&E-CSE-G PM Mr. Vann, S&E-CSE-GA Mr. Grau, S_E-QUAL-DIR Mr. Han_ners, S&E-CSE-I Mr. Chandler, S&E-QUAL-DIR Mr. L. James, PM-DIR Mr. Wolfe, S&E-CSE-! Mr. Henritze, S&E-QUAL-A Mr. Andressen, PM-PR-CM Mr. E. May, S&E-CSE-L Mr. Rushing, S&E-QUAL-PI Col. Teir, PM-SAT-IB-MGR Mr. McKay, S&E-CSE-LF Mr. Klauss, S&E-QUAL-J Mr. Huff, PM-SAT-E Mr. R. L. Smith, S&E-CSE-V Mr. Hughes, S&E-QUAL-P Dr. Speer, PM-MO-MGR (4) Mr. Brooks, S&E-CSE-V Mr. Landers, S&E-QUAL-PC (3) Mr_ Belew, PM-AA-MGR Mr. Hagood, S&E-CSE-M (3) Mr. Peck, S&E-QUAL-F Mr. Brown, PM-EP-MGR Mr. Brien, S&E-QUAL-Q Mr. Smith, PM-EP-J S&E-ASTR Mr. Wittmann, S&E-QUAL-T V. J. Norman, PM-MO Mr. Davis, S&E-QUAL-F Mr. Stewart, PM-EP-F Mr. Moore, S&E-ASTR-DIR Mr. Godfrey, PM-SAT-MGR Mr. Stroud, S&E-ASTR-SC S&E-SSL Mr. R. Smith, PM-SAT-MGR Mr. Robinson, S&E-P-ATM (4487) Mr. Burns, PM-SAT-T Mr. Erickson, S&E-ASTR-SE Mr. Heller, S&E-SSL-DIR Mr. Bell, PM-SAT-E Mr. Darden, S&E-ASTR-SDC Mr. Sieber, S&E-SSL-S Mr. Rowan, PM-SAT-E Mr. Justice, S&E-ASTR-SDA Mr. Moody, PM-SAT-Q Mr. Vallely, S&E-ASTR-SDC MS Mr. Webb, PM-SAT-P Mr. George, S&E-ASTR-SDI Mr. Urlaub, PM-SAT-S-IB/S-IC Mr. Mandel, S&E-ASTR-G MS-H Mr. Lahatte, PM-SAT-S-II Mr. Ferrell, S&E-ASTR-GS MS-I Mr. McCullough, PM-:SAT-S-IVB Mr. Powell, S&E-ASTR-I MS-IP Mr. Duerr, PM-SAT-IU Mr. Avery, S&E-ASTR-SC MS-IL (8) Mr. Smith, PM-SAT-G Mr. Kerr, S&E-ASTR-IRD MS-D Col. Montgomery, PM-KM Mr. Threlkeld, S&E-ASTR-ITA Mr. Peters, PM-SAT-S-IVB Mr. Boehm, S&E-ASTR-M CC-P Mr. Weir, PM-SAT-IU Mr. Lominick, S&E-ASTR-GMF Mr. Ferrell, PM-EP-EJ Mr. Taylor, S&E-ASTR-R Mr. Wofford, CC-P Dr. Constan, PM-MA-MGR Mr. Riemer, PM-MA-QP S&E COMP Mr. Balch, PM-MT-MGR KSC Mr. Auter, PM-MT-T Dr. Hoelzer, S&E-COMP-DIR Mr. Sparks, PM-SAT-G Mr. Prince, S&E-COMP-DIR Dr. Debus, CD Mr. Ginn, PM-SAT-E Mr. Fortenberry, S&E-COMP-A Adm. Middleton, AP (5) Mr. Hal ey, PM-SAT-S-IB/S-IC Mr. Cochran, S&E-COMP-R Dr. Gruene, LV Mr. Higgins, PM-SAT-S-IVB Mr. Houston, S&E-COMP-RR Mr. Rigell, LV-ENG Mr. Odom, PM-SAT-S-II Mr. Craft, S&E-COMP-RR Mr. Sendler, IN Mr. Stover, PM-SAT-S-II Mr. Mathews, AP Mr. Reaves, PM-SAT-Q S&E-ME Dr. Knothe, EX-SCl Mr. Wheeler, PM-EP-F Mr. Edwards, LV-INS Mr. Johnson, PM-SAT-T Mr. Siebel, S&E-ME-DIR Mr. Fannin, LV-MEC Mr. Cushman, PM-SAT-T (10) Mr. Wuencher, S&E-ME-DIR Mr. Pickett, LV-TMO Mr. Marchese, PM-MA-QR Mr. Orr, S&E-ME-A Mr. Rainwater, LV-TMO Mr. Franklin, S&E-ME-T Mr. Bell, LV-TMO-3 S&E Mr. Lealman, LV-GDC S&E-ASTN Mr. Preston, DE Mr. Weidner, S&E-DIR Mr. Mizell, LV-PLN-12 Mr. Richard, S&E-DIR Mr. Heimburg, S&E-ASTN-DIR Mr. O'Hara, LV-TMO Dr. Johnson, S&E-R Mr. Kingsbury, S&E-ASTN-DIR Mr. Brown, AP-SVO-3 Mr. Hamilton, MSC-RL Mr. Smith, AP-SVO EXTERNAL

Headquarters, National Aeronautics & Space Administration Office of the Asst. Sec. of Defense for Research Washington, D. C. 20546 and Engineering Room 3EI065 Dr. Mueller, M The Pentagon Mr. Petrone, MO Washington, D.C. 20301 Gen. Stevenson, MO (3 copies) Attn: Tech Library Mr. Hage, MO Mr. Schneider, MO-2 Director of Guided Missiles Capt. Freitag, MC Office of the Secretary of Defense Capt. Holcomb, MAO Room 3E131 Mr. White, MAR (2 copies) The Pentagon Mr. Day, MAT (I0 copies) Washington, 0. C. 20301 Mr. Wilkinson, MAB Mr. Kubat, MAP Central Intelligence Agency Mr. Wagner, MAS (2 copies) Washington, D.C. 20505 Mr. Armstrong, MB Attn: OCR/DD/Publications (5 copies) Mr. Mathews, ML (3 copies) Mr. Lord, MT Director, National Security Agency Mr. Lederer, MY Ft. George Mead, Maryland 20755 Attn: C3/TDL Director, Ames Research Center: Dr. H. Julian Allen National Aeronautics & Space Administration U. S. Atomic Energy Commission, Sandia Corp. Moffett Field, California 94035 University of California Radiation Lab. Technical Information Division Director, Flight Research Center: Paul F. Bikle P. O. Box 808 National Aeronautics & Space Administration Livermore, California 94551 P. O. Box 273 Attn: Clovis Craig Edwards, California 93523

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4 EXTERNAL (CONT.)

J. E. Trader NASA Resident Manager's Office McDonnell Douglas Astronautics Corp. 5301 Balsa Avenue Huntington Beach, California 92646

L. C. Curran NASA Resident Manager's Office North American Rockwell/Space Division 12214 Lakewood Blvd. Downey, California 90241

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W. Klabunde Northrup 6025 Technology Drive Huntsville, Alabama 35804

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