Saturn V Launch Vehicle Flight Evaluation Report - AS-505 Apollo 10 Mission

8/7/2019 Saturn V Launch Vehicle Flight Evaluation Report - AS-505 Apollo 10 Mission

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GEORGEC. MARSHALL SPACE FLIGHTCENTER

MPR-SAT-FE-9-7

SATURNV LAUNCHVEHICLE _

FLIGHT I'VALUATIONREPORT-AS-50_ ,,i

,A P O LLO 10 M I S S IO N

PREPAREDBY

SATURNFLIGHTEVALUATIONWORKINGGROUP



C &0 r_ Z I c mI imm



MP R-S A T -F E -69 -7

S A T UR N V L A U N C H V E HIC L E F L IG H T E V A L UA T IO N R EP O R T - A S -50 5

A P O LL O I0 M IS S IO N

B Y

S aturn F light E valuation W orking Group

G eorge C. M arsha ll S pace F ligh t C enter

A B S T R A C T

S atu rn V A S -50 5 (A pollo lO M ission) was launched at 1 2:49 :0 0 E astern

Dayligh t T im e on M ay 1 8, 1 9 6 9 , from Kennedy S pace C enter, C om plex 39 ,P ad B . T he veh ic le lifted o ff on schedule on a launch azim uth o f 9 0

degrees east of north and ro lled to a fli.gh t azim uth o f 7 2.0 28 degreeseast of north .

T he la unch veh icle successfu lly p laced th e m an ned spacecra ft in the

planned translunar injection coast mode. The S-IVB/IU was placed in aso la r o rbit w ith a period o f 344.9 days by a com bination o f con tinuous

L H 2 vent, the con tingency experim ent of propellan t lead, a L O X dum p and

A P S ullage burn.

T he M ajor F light O bjectiw_s and the Deta iled T est O bjectives of th is

m ission were com plete ly accom plished. N o fa ilu res, anom alies, or de-

via tions occurred that seriously affected the flight or m ission.

Any questions or comments pertaining to the information contained inthis report are invited arid should be directed to:

Director, G eorge C. M arsha ll S pace F ligh t Center

Huntsville_, A labama 3581 2

A ttention: Chairm an, S aturn F ligh t E va luation W orking

G roup, S &E -C S E -L F (P hon e 453-257 5)



TABLE OF CONTENTS

Section Page

TABLEOFCONTENTS iii

LISTOFILLUSTRATIONS xii

LISTOFTABLES xix

ACKNOWLEDGEMENT xxiii

ABBREVIATIONS xxiv

MISSIONLAN xxvii

FLIGHTESTSUMMARY xxix

1 INTRODUCTION

l.l Purpose I-I

1.2 Scope I-I

2 EVENTTIMES

2,I Summaryof Events 2-I

2.2 V ariable T im e and C om m andedS witch

SelectorEvents 2-3

3 L A U N C HO P E R A T IO N S

3.1 Summary 3-I

3.2 PrelaunchMilestones 3-I

3.3 Countdownvents 3-I

3.4 Propellantoading 3-I

3.4.1 RP-ILoading 3-I

3.4.2 LOXLoading 3-3

3.4.3 LH2Loading 3-3

3.4.4 Auxiliary Propulsion System

Propellantoading 3-3

3.5 S-II Insulation, Purge and LeakDetection 3-3

3.5.1 ForwardBulkheadInsulation 3-3

3.5.2 Forward Bulkhead Uninsulated A rea 3-4

3.5.3 LH2 TankSidewall 3-4

ii i



T AB L E O F C O N TE N T S (C O N TIN UE D)

Section Page

3.5.4 Bolting/J-Ring 3-43.5.5 FeedlineElbow 3-4

3,5.6 Commonulkhead 3-4

3,6 GroundSupport Equipment 3-5

3.6.1 Ground/Vehicle Interface 3-5

3.6 .2 M S F C F urn ished G round S upportEquipment 3-5

3.6.3 GSECameraCoverage 3-6

4 TRAJECTORY

4.1 Summary 4-I

4.2 Tracking Data Utilization 4-I

4.2.1 T rack ing D uring the A scent P hase

of Flight 4-I4.2.2 Tracking During Orbital Flight 4-2

4.2.3 T racking D uring the In jection P hase

of Flight 4-2

4.3 Trajectory Evaluation 4-2

4.3.1 Ascent Trajectory 4-2

4.3.2 Parking Orbit Trajectory 4-3

4.3.3 Injection Trajectory 4-5

4.3.4 Post TLI Trajectory 4-6

4.3.5 S-IVB/IU Post Separation Trajectory 4-10

5 S-IC PROPULSION

5.1 Summary 5-I

5.2 S-IC Ignition Transient Performance 5-I

5.3 S-IC Mainstage Performance 5-4

5 .4 S - IC E n g in e S h u tdo w n T ra n s ien tPerformance 5-6

5,5 S-IC Stage Propellant Management 5-6

5.6 S-IC Pressurization Systems 5-6

5.6.1 S-IC Fuel Pressurization System 5-6

5.6.2 S-IC LOXPressurization System 5-7

5.7 S-IC Pneumatic Control Pressure System 5-85.8 S-ICPurgeSystems 5-9

5.9 POGOSuppression System 5-9

iv



T A B L E O F C O N T E N TS (C O NT IN U E D)

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 6-5

6.4 S-II Shutdown Transient Performance 6-7

6.5 S-II Stage Propellant Management 6-9

6.6 S-II Pressurization Systems 6-10

6.6.1 S-II Fuel Pressurization System 6-10

6.6.2 S-II LOXPressurization System 6-11

6 .7 S - I I P n eu m a tic C o n tro l P ressu reSystem 6-11

6.8 S-II Helium Injection System 6-12

7 S-IVB PROPULSION

7.1 Summary 7-I

7 .2 S - IV B C h illdow n a n d B u ildu p T ran sien tPerformance for First Burn 7-2

7 .3 S - IV B M a ins ta ge P e rfo rm a nce fo r F irs tBurn 7-2

7.4 S-IVB Shutdown Transient Performancefor First Burn 7_5

7.5 S-IVB Parking Orbit Coast Phase

Conditioning 7-6

7.6 S-IVB Chilldown and Restart for

SecondBurn 7-6

7 .7 S - IV B M a in s ta g e P erfo rm a n ce fo rSecondBurn 7-12

7 .8 S -IV B S hu tdow n T ransie n t P erfo rm ance

for' SecondBurn 7-14

7.9 S-IVB Stage Propellant Management 7-14

7.10 S-IVB Pressurization System 7-14

7.10.1 S-lVB Fuel Pressurization System 7-14

7.10.2 S-IVB LOX Pressurization System 7-18




Section Page

7 .1 1 S -IV B P neum atic C ontro l P ressu re

System 7-22

7.12 S-IVB Auxiliary Propulsion System 7-26

7 .1 3 S -IV B P rope llan t L ead E xperim ent and

Orbital Safing Operation 7-27

7 .1 3,1 L O X an d L H2 L ea d C h illdo w n E xpe rim e n t 7 -287.13.2 LOX Tank Ambient Repressurization 7-31

7.13.3 LH2 Tank Ambient Repressurization 7-327.13.4 Fuel TankSafing 7-32

7.13.5 LOXTank Dumpand Safing 7-33

7.13.6 Cold HeliumDump 7-33

7.13.7 Ambient Helium Dump 7-34

7 .1 3.8 S tage P neum atic C ontro l S phere

Safing 7-35

7.13.9 Engine Start Sphere Safing 7-35

7.13.10 Engine Control Sphere Safing 7-35

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 (First Burn) 8-I

8.5 S - IV B H ydra u lic S ystem (P a rk in g O rb itCoastPhase) 8-2

8.6 S-IVB Hydraulic System (Second Burn) 8-3

8,7 S - IV B H ydra u lic S ystem (T ra n s lu n a rIn je c tio n C o as t an d P ro pe lla n t D um p ) 8 -3

9 STRUCTURES

9,1 Summary 9-I

9.2 Total Vehicle Structures Evaluation 9-I

9.2.1 Longitudinal Loads 9-I

9,2.2 BendingMoments 9-3

9.2.3 Vehicle Dynamic Characteristics 9-5

9.3 Vibration Evaluation 9-11

9.3.1 S-IC Stage and Engine Evaluation 9-11

9.3.2 S-II Stage and Engine Evaluation 9-12

9.3.3 S-IVB Stage and Engine Evaluation 9-23

vi



T A B LE O F C O N TE N T S (C O N TIN UE D)

Section Page

I0 GUIDANCEAND NAVIGATION

ID.1 Summay 1O-1

I0.I.I Flight Program I0-I0.1.2 Instrument Unit Components I0-I

10.2 GuidanceComparisons I0-I

10.3 Navigation and Guidance SchemeEvaluation 10-7

1 0 .4 G u ida n ce S ystem C o m p o n en t E va lu a tio n 1 0 -810.4.1 LVDCPerformance 10-8

10.4.2 LVDAPerformance 10-8

10.4.3 LadderOutputs 10-810.4.4 Telemetry Outputs I0-11

10.4.5 Discrete Outputs I0-II10.4.6 Switch Selector Functions I0-II1 0 .4.7 S T -1 24M -3 Inertia l P latform

Performance 10-I1

11 CONTROLSYSTEM

11.1 Summary 11I

II.2 Control System Description II-I

11.3 S-IC Control System Evaluation 11-211.3.1 Li ftoff Clearances 11-2

11.3.2 S-IC Flight Dynamics II-2

II.4 S-II Control System Evaluation II-7

II.5 S-IVB Control System Evaluation 11-14

1 1 .5.1 C ontro l S ystem E va lua tion D uringFirstBurn 11-14

II.5.2 Control System Evaluation During

Parkingrbit ll-14

'II.5.3 Control System Evaluation DuringSecondBurn ll-17

11.5.4 Control System Evaluation After

S-IVBSecondBurn ll-18

12 SEPARATION

12.1 Summary 12-I

_\ 12.2 S-IC/S-II Separation Evaluation 12-I

12.2.1 S-IC RetroMotorPerformance 12-1

12.2.2 S-IllUllageMotor Performance 12-I

12.2.3 S-IC/S-II Stage Separation 12-1

_\, vio _




Section Page

1 2.3 S - II S e co nd P la n e S e pa ra tio nEvaluation 12-I

12.4 S-II/S-IVB Separation Evaluation 12-212.4.1 S-II Retro Motor Performance 12-2

12.4,2 S-IVB Ullage Motor Performance 12-2

12.4.3 S-II/S-IVB Separation Dynamics 12-2

12.5 S-I VB/IU/LM/CSM SeparationEvaluation 12-2

1 2.6 L un a r M odu le D ock ing a n d E je ctio nEvaluation 12-2

1 3 E L E C T R IC A L N E T W O R K S

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

1 4 R A N G E S A F E T Y A N D C O M M A N DS Y S T E M S

14.1 Summary 14-I

14.2 Secure Range Safety CommandSystems 14-I

14.3 Commandand Communications System 14-I

1 5 E M E R G E N C YD E T E C T IO N S Y S T E M

15.1 Summary 15-I

15.2 SystemEvaluation 15-I15.2.1 General Performance 15-I

15.2.2 Propulsion System Sensors 15-I

1 5 ,2.3 F lig h t D yn a m ics a n d C o n tro l S en so rs 1 5 - I15.2.4 EDSEventTimes 15-2

1 6 V E H IC L E P R E S S U R E A N D A C O U S T IC E N V IR O N M E N T

16.1 Summary 16-I

1 6 .2 S urface P ressu res and C om partm en tVenting 16-I

16.2.1 S-ICStage 16-I

16.2.2 S-II Stage 16-5

viii




Section Page

16.3 BasePressures 16-5

16.3.1 S-lC BasePressures 16-5

16.3.2 S-II BasePressures 16-7

16.4 Acoustic Environment 16-816.4,1 External Acoustics 16-8

16.4.2 Internal Acoustics 16-17

1 7 V I--H IC L E T H E R M A L E N V IR O N M E N T

17.1 Summary 17-I

17.2 S-IC Base Heating and StageSeparation Environment 17-I

17.2.1 S-IC BaseHeating 17-I17.2.2 S-IC/S-II Separation Environment 17-3

17.3 S-II Base Region Environment 17-3

17,4 Vehicle Aeroheating ThermalEnvironment 17-I0

1 7 .4.1 S -IC S tag e A e ro he a tin g E n v iron m en t 1 7 -1 0

1 7 .4.2 S -II S tag e A e ro he a tin g E n v iron m en t 1 7 -1 2

1 7 .4.3 S -IV B S tag e A e rohea ting E nv ironm e n t 1 7 -1 7

1 8 E N V IR O N M E N T A L C O N T R O LS Y S T E M

18.1 Summary 18-I

18.2 S-IC Environmental Control 18-I

18.3 S-II Environmental Control 18-5

18.4 IU Environmental Control 18-5

18.4.1 Thermal Conditioning System 18-6

18.4.2 Gas Bearing Supply System 18-8

19 DATASYSTEMS

19.1 Summary 19-1

19.2 Vehicle Measurements Evaluation 19-I

I!).3 Airborne Telemetry Systems 19-5

I!).4 Airborne Tape Recorders 19-5

19.5 RFSystemsEvaluation 19-71 9 .5.1 T elem etry S ystem R F P ropagation

Evaluation 19-7

1 9 .5.2 T rack ing S ystem s R F P ropagationEvaluation 19-8

ix



T AB LE O F C O N TE N T S (C O N TIN UE D)

Section Page

19.5.3 CommandSystems RFEvaluation 19-9

19.6 Optical Instrumentation 19-13

20 MASS CHARACTERISTICS

20.1 Summary 20-I

20.2 MassEvaluation 20-I

21 MISSIONOBJECTIVESACCOMPLISHMENT 21-I

22 F A IL U R E S , A N O M A L IE S A N D D E V IA T IO N S

22.1 Summary 22-I

22.2 System Failures and Anomalies 22-I

22.3 SystemDeviation 22-I

23 SPACECRAFTUMMARY 23-I

Appendix

A ATMOSPHERE

A.I Summary A-I

A.2 General Atmospheric Conditions atLaunchTime A-I

A.3 Surface Observations at Launch Time A-I

A.4 UpperAir Measurements A-2

A.4.1 WindSpeed A-2

A.4.2 WindDirection A-3

A.4.3 Pitch WindComponent A-3

A.4.4 YawWindComponent A-7A.4.5 ComponentWindShears A-7

A .4 .6 E xtre m e W ind D a ta in th e H ig h D yn a m ic

PressureRegion A-7

A.5 Thermodynamicata A-7A.5.1 Temperature A-7

A.5.2 Atmospheric Pressure A-7A.5.3 Atmospheric Density A-13

A.5.4 Optical Index of Refraction A-13

A .6 C o m parison o f S e lecte d A tm osphe ricData for Saturn V Launches A-13



T AB L E O F C O N T EN T S (C O N T IN UE D)

Appendix Page

B A S -50 5 S IG N IF IC A N T C O N F IG U R A T IO N C H A N G E S

B.I Introduction B-I

x i





L IS T O F ]IL LUS TR AT IO N S (CO N TIN UE D)

Figure Page

6-8 S-II LOXPumpnlet Conditions 6-15

7 - I S - IV B S ta rt B o x a n d R u n R equ irem en ts - F irs t B u rn 7 -3

7-2 S-IVB Steady State Performance - First Burn 7-4

7-3 S-IVB CVSPerformance- Coast Phase 7-7

7 -4 S -IVB U llage P ressure During R epressuriza tion

Using 02/H2 Burner 7-9

7-5 S-IVB 02/H2 Burner Thrust and Pressurant Flowrate 7-10

7 -6 S -IV B S ta rt B o x a n d R u n R e qu ire m en ts - S eco n dBurn 7-11

7-7 S-IVB Steady State Performance - Second Burn 7-13

7-8 S-IVB LH2 Ullage Pressure- First Burn and

Parkingrbit 7-16

7-9 S-IVB LH2 Ullage Pressure - Second Burn andTranslunaroast 7-17

7-10 S-IVB F uel P ump Inlet Conditions- First B urn 7-19

7-11 S-IVB F uel Pump Inlet C onditions - Second Burn 7-20

7-12 S-IVB LOX T ank Ullage Pressure - First B urn

andParkingOrbit 7-21

7-13 S-IVB LOXTank Ullage Pressure - Second Burn,TranslunarCoast 7-22

7-14 S-IVB LOX PumpInlet Conditions - First Burn 7-23

7 -1 5 S -IV B L O X P u m p In le t C o n d itio n s - S eco n d B u rn 7 -24

7-16 S-IVB ColdHelium Supply History 7-25

7-17 S-IVB APSHeliumBottle Mass 7-26

7-18 S-IVB APSHelium Bottle Temperature 7-28

7 -1 9 S -IVB A P S P ropellan ts R em ain ing Versus R angeTime, ModuleNo. 1 andModuleNo. 2 7-29

7 -20 S -IVB P ropellant L ead E xperim ent and O rbita lSafingSequence 7-32

7-21 LOXPumpInlet Chilldown Effectiveness 7-34

7 -2 2 L O X P u m p Disch a rg e C h illdo w n E ffec tiv en ess 7 -35

7-23 S-IVB Fuel Lead Chilldown Effectiveness 7-36

7-24 S-IVBLOXDump 7-37

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L IS T O F IL LUS TR AT IO N S (CO N TIN UE D)

Figure Page

8-I S-ICEngine,ValveClosingPressure 8-2

8-2 S-IVB HydraulicSystem Pressure- Second Burn 8-4

8-3 S-IVB Hydraulic S ystem Actuator Position -Secondurn 8-5

8-4 S-IVB Auxiliary Hydraulic Pump Performance -3econdurn 8-6

8-5 S-IVB Auxiliary Hydraulic Pump Performance -CoastPhaseandThirdThermalCycle 8-7

8-6 S-IVB Hydraulic S ystem Actuator Positions -CoastPhaseandThirdThermalCycle 8-8

8-7 S-IVB Auxiliary Hydraulic Pump Performance -CoastPhaseandPassivation 8-9

8-8 S-IVB Hydraulic System Actuator Positions -CoastPhaseandPassivation 8-I0

9-I ReleaseRodForce- Displacementurves 9-2

9-2 Longitudinal Loads at Maximum Bending Moment,

Center Engine Cutoff, and Outboard Engine Cutoff 9-3

9-3 Longitudinal S tructural Dynamic Response DuetoOutboardngineCutoff 9-4

9-4 MaximumBendingMomentNearMaxQ 9-5

9-5 First Longitudinal M odal Frequencies During

S-ICPoweredlight 9-6

9-6 Peak Amplitudes of Vehicle First LongitudinalModeforAS-504andAS-505 9-7

9-7 Comparison of AS-504 and AS -505 S-II S tageLowFrequencyscillations 9-9

9-8 S-IVBSecondBurn45 HertzOscillations 9-I0

9-9 AS-505 LateralAnalysis/Measuredodal

Frequencyorrelation 9-12

9-I0 S-ICStageStructureVibrationEnvelopes 9-14

9-11 S-ICStageEngineVibrationEnvelopes 9-15

9-12 S-IC Stage ComponentsVibrationEnvelopes 9-16

9-13 S-ICVibrationMeasurementLocations 9-17

9-14 S-IIStageStructureVibrationEnvelopes 9-19

x i v




Figure Page

9-15 S.-II Stage Engine Vibration Envelopes 9-20

9-16 S--II Stage Component Vibration Envelopes 9-219-17 S--IVBStage Vibration Envelopes 9-24

9-18 S--IVB Stage Engine Vibration Envelopes 9-25

I0 -I T racking and S T -1 24M -3 P la tfo rm VelocityComparison (Trajectory Minus Guidance) 10-2

1 0 -2 A ttitude C om m ands During A ctive G u idance P eriod 1 0 -9

10-3 Orbital Attitude Commands I0-I0

II-I Pitch Plane DynamicsDuring S-IC Burn 11-4

11-2 YawPlane DynamicsDuring S-IC Burn 11-5

11-3 Roll Plane DynamicsDuring S-IC Burn 11-6

11-4 Normal Acceleration During S-IC Burn 11-8

1 1 -5 P itch and Y aw P lane W ind Velocity and F ree-S treamAngle-of-Attack During S-IC Burn 11-9

II-6 Pitch Plane DynamicsDuring S-II Burn II-II

11-7 YawPlane DynamicsDuring S-II Burn 11-12

11-8 Roll Plane DynamicsDuring S-II Burn 11-13

1 1 -9 P itch A ttitude C o ntro l D u ring S -IV B F irs t B u rn 1 1 -1 5

I I- I0 Y a w A ttitu de C o n tro l Du rin g S -IV B F irst B u rn 1 1 -1 5

II-II Roll Attitude Control During S-IVB First Burn 11-16

1 1 -1 2 P itch A ttitu de C o n tro l D u rin g P a rk in g O rbit 1 1 -1 7

1 1 -1 3 P itch A ttitude C ontro l During S -IVB S econd B urn 1 1 -1 9

1 1 -1 4 Y aw A ttitu de C on tro l D urin g S -IV B S econd B urn 1 1 -20

1 1 -1 5 R o ll A ttitude C ontro l D uring S -IV B S econd B urn 1 1 -20

1 1 -1 6 P itch A ttitude C ontro l A fter S -IVB S econd B urn 1 1 -21

1 3-I S -- IVB S tage F orward B a ttery N o. 1 Vo ltageandCurrent 13-4

1 3-2 S -IVB S tage F orward B a ttery N o. 2 Vo ltageandCurrent 13-4

1 3-3 S -.IVB S tage A ft B a ttery N o. 1 V oltage andCurrent 13-5

13-4 S-IVB Stage Aft Battery No. 2 Voltage andCurrent 13-5

×V



L IS T O F ILL US TR AT IO N S (CO N TIN UE D)

Figure Page

1 3-5 B a tte ry 6 D IO V o ltag e, C u rren t and T em pera tu re 1 3-7

1 3-6 B a tte ry 6 D 30 V o ltag e, C u rren t and T em pera tu re 1 3-8

1 3-7 B a tte ry 6 D 40 V o ltag e, C urren t and T em pera tu re 1 3-9

1 6 -I S -IC E ng ine F a iring Com partm ent P ressureDifferential 16-2

1 6 -2 S - IC C o m p a rtm e n t P ressu re D iffe ren tia ls 1 6 -3

16-3 S-IC CompartmentPressure Loading 16-4

16-4 S-II Forward Skirt Pressure Loading 16-6

16-5 S-IC BasePressure Differential 16-7

16-6 S-IC BaseHeat Shield Pressure Loading 16-8

1 6 -7 T hrust C one and B a se H eat S h ie ld F orw ardFacePressures 16-9

16-8 S-II Heat Shield Aft FacePressures 16-10

1 6 -9 Veh icle E xterna l O vera ll S ound P ressure L evelat Liftoff 16-11

1 6 -1 0 Veh icle E xternal S ound P ressure S pectra lDensities at Liftoff 16-12

1 6 -1 1 Vehicle E xterna l O vera ll F luctuating P ressureLevel 16-14

1 6 -1 2 Vehicle E xterna l F luctuating P ressure S pectra lDensities at M aximum Inflight A erodynamic N oise 1 6-1 6

1 6 -1 3 S -IC Heat S h ie ld P anels In terna l A cousticEnvironment 16-17

1 6 -1 4 S -IC Intertank Internal A coustic E nvironm ent 1 6 -1 8

16-15 S-II Internal Acoustics History 16-19

1 6 -1 6 S - IV B A co u stics L ev e ls D u rin g S -IC B u rn 1 6 -20

1 7 - I S - IC B a se H ea t S h ie ld T h erm a l E n v iro n m en t 1 7 -2

17-2 F-I EngineThermal Environment 17-2

1 7 -3 S -IC H ea t S h ie ld F o rw a rd S u rfa ce T em p era tu re 1 7 -417-4 S-IC Heat Shield Bondline Temperature 17-4

1 7 -5 S - IC B a se H e a t S h ie ld M ea su rem en t L o ca tio n s 1 7 -5

1 7 -6 S -IC T em perature Under Insu la tion , Inboard S ideEngineNo.1 17-6

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L IS T O F IL LUS TR AT IO NS (C ON TIN UE D)

Figure Page

17-7 S-IC Forward Skirt Compartment Ambient Air

TemperatureDuringS-IC/S-IIStage Separation 17-6

17-8 S-II Heat S hield Base R egion Heating Rates 17-717-9 S-IIThrustConeHeatingRate 17-8

17-10 S-IIBaseGasTemperature 17-8

17-11 S-IIHeatShieldAftFaceTemperatures 17-9

17-12 S-ICIntertankAerodynamicHeating 17-11

17-13 S-IC Forward S kirt A erodynamicHeating 17-11

17-14 ForwardLocationof SeparatedFlow 17-12

17-15 S-ICLOXTankSkinTemperature 17-13

17-16 S-ICFuelTankSkinTemperature 17-13

17-17 S-ICIntertankSkinTemperatures 17-14

17-18 S-ICForwardSkirtSkinTemperature 17-14

17-19 S-II Aft Interstage A eroheating Environment 17-15

17-20 S-II Aft Interstage Aeroheating Environment,

UllageotorFairing 17-16

17-21 S-II Aft Interstage Aeroheating Environment,

LH2 FeedlineftFairing 17-17

17-22 S-II Body Aeroheating Environment, LH2 Feedline

Forwardairing 17-18

17-23 S-II Body Aeroheating Environment, ForwardSkit 17-I

17-24 S-II Body Aeroheating Environment, Systems

TunnelForwardairing 17-19

18-1 S-IC Forward Compartment Canister Temperature 18-2

18-2 S-IC Forward Compartment Ambient Temperature 18-3

18-3 S-ICAft CompartmentTemperatureRange 18-4

18-4 RTGPurgeDuctingModification 18-6

18-5 IU SublimatorPerformanceDuringAscent 18-7

18-6 Thermal Conditioning System GN 2 Sphere Pressure(D25-601) 18-9

18-7 Flight Control Computer Temperatures (C69-602) 18-9

18-8 SelectedComponentTemperatures 18-10

x v i i




Figure Page

18-9 GasBearing GN2 SpherePressure 18-11

18-10 Inertial Platform GN Pressures 18-11

19-I VHFTelemetry Coverage'Summary 19-8

19-2 C-BandRadarCoverageSummary 19-10

19-3 CCSSignal Strength Fluctuations at Guaymas 19-11

1 9 -4 C C S S igna l S trength F luctua tions a t G o ldstone 1 9 -1 2

19-5 CCSSystemBlock Diagram 19-14

1 9 -6 E lec tr ica l S ch em a tic o f C C S C o a xia l S w itch 1 9 -1 4

19-7 CCSCoverageSummary 19-15

A-I Scalar WindSpeed at Launch Time of AS-505 A-4

A-2 Wind Direction at LaunchTime of AS-505 A-5A -3 P itch W in d S p eed C o m p on ent (Wx) a t L au n ch

TimeofAS-505 A-6

A -4 Y a w W in d S p eed C o m p o n en t a t L a u n ch T im eofAS-505 A-8

A-5 Pitch (Sx) and Yaw (Sz) ComponentWihd Shearsat LaunchTimeof AS-505 A-9

A -6 R ela tive Devia tion of T em peratu re and DensityF rom the P R A -6 3 R eference A tm osphere, A S -50 5 A -II

A -7 R ela tive D ev ia tion o f P ressure and A bso lute

Deviation of the Index of R efraction F rom thePPJ_-63Reference Atmosphere, AS-505 A-12

x v i i i



L IS T O F T A B LE S

Table Page

2-I TimeBaseSummary 2-3

2-2 SignificantEvent Times Summary 2-4

2-3 Variable Time and Commanded Switch Selector

Events 2-I0

3-I AS-505Prelaunchilestones 3-2

4-I Comparisonof SignificantTrajectoryEvents 4-7

4-2 ComparisonfCutoffEvents 4-8

4-3 Comparisonof Separation Events 4-9

4-4 StageImpactLocation 4-10

4-5 Parking Orbit Insertion Conditions 4-11

4-6 Translunar Injection Conditions 4-14

4-7 Comparisonof Slingshot Maneuver 4-14

4-8 Lunar CloseApproachParameters 4-16

4-9 Heliocentric Orbit Parameters 4-16

5-I S-IC Engine PerformanceDeviations 5-4

5-2 S--ICStage Propellant MassHistory 5-7

6 -I S --II E ng ine P erfo rm ance Devia tions(ESC+61Seconds) 6-7

6-2 S.-II Propellant MassHistory 6-10

7 -I S -- IV B S teady S ta te P erfo rm an ce - F irs t B urn(E S C +1 40 -S econd T ime S lice at S tandardAltitude Conditions) 7-5

7 -2 S ,- IVB S teady S ta te P erfo rm ance - S econd B urn(E S C +lS O -S econd T ime S lice at S tandardAltitude Conditions) 7-15

7-3 S,-IVBStage Propellant MassHistory 7-15

7-4 S-IVB APSPropellant Conditions 7-27

7-5 S-IVB APSPropellant Consumption 7-30

xix



L IS T O F T A B LE S (C O N T IN UE D )

Tabe Page

7-6 S-IVB HeliumBottle Conditions 7-31

9-I S-IVB Stage Low Frequency Vibration Summary 9-11

9-2 S-IC Stage Vibration Summary 9-13

9-3 S-II Stage MaximumOverall Vibration Levels 9-18

9-4 S-IVB Vibration Summary 9-26

I0-I Inertial Platform Velocity Comparisons 10-3

10-2 GuidanceComparisons 10-5

10-3 Guidance ComponentComparisons 10-7

10-4 Start and Stop Times for IGMGuidance Commands 10-8

10-5 Translunar Injection Parameters I0-II

I I - I A S -50 5 M isa lignm en t and L ifto ff C on ditions

Summary 11-3

1 1 -2 M a x im um C o n tro l P a ra m e ters D u rin g S -IC B o o stFlight 11-7

1 1 -3 M a x im um C o n tro l P a ra m e ters D u rin g S -II B o o st

Flight II-I0

11-4 Maximum Control Parameters During S-IVBFirst Burn 11-16

1 1 -5 M a x im um C o n tro l P a ra m e ters D u rin g S -IV B

SecondBurn 11-1813-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 PowerConsumption 13-6

1 4-I C om m and and C om m un ica tion s S ystem C om m an dsHistory, AS-505 14-2

15-I MaximumngularRates 15-I

15-2 EDSRelatedEvent Times 15-2

15-3 EDSAssociated Discretes 15-3

16-I Sound Pressure Level Comparison of AS-505With Previous Saturn V Flight Data 16-19

18-I TCSCoolant Flowrates and Pressures 18-7

19-I AS-505 Flight MeasurementSummary 19-2

XX



L IS T O F T A B L E S (C O N T IN UE D)

Tabe Page

1 9 -2 A S -5 0 5 F lig h t M ea su re m en ts W a iv e d P rio r toLaunch 19-2

19-3 AS-505MeasurementMalfunctions 19-3

19-4 AS-505 Questionable Flight Measurements 19-5

19-5 AS-505 Launch Vehicle Telemetry Links 19-6

19-6 TapeRecorderSummary 19-6

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 -

Ki1ograms 20-7

20-6 Total Vehicle Mass - S-IVB First Burn Phase -PoundsMass 20-8

20-7 Total Vehicle Mass - S-IVB Second Burn Phase -

Ki1ograms 20-9

20-8 Total Vehicle Mass - S-IVB Second Burn Phase -

PoundsMass 20-10

20-9 Flight SequenceMassSummary 20-11

20-I0 Mass Characteristics Comparison 20-13

21-I Mission Objectives Accomplishment Summary 21-I

22-I H ardw are C ritica lity C a tego rie s fo r F lig h tHardware 22-I

22-2 Summaryof Failures and Anomalies 22-2

22-3 Summaryf Deviations 22-2

A-I Surface Observations at AS-505 Launch Time A-I

A-2 Solar Radiation at AS-505 Launch Time, LaunchPad39B A-2

A-3 Systems Used to Measure Upper Air Wind Datafor AS-505 A-3

A-4 MaximumWind Speed in High Dynamic Pressure

Region for Apollo/Saturn 501 through Apollo/Saturn505Vehicles A-IO

x x i



L IS T O F T A BL E S (C O N T IN UE D)

Tabe Page

A-5 Extreme Wind Shear Values in the High Dynamic

Pressure Region for Apollo/Saturn 501 through

Apollo/Saturn 505 Vehicles A-IO

A -6 S e le cted A tm ospheric O bserva tions fo r A p o llo /

Saturn 501 through Apollo/Saturn 505 VehicleLaunches 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

x x i i



A CKN OWLE DG E ME N T

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 E ngineering

Central Systems EngineeringA er o-A str ody na mi cs L abora to ry

A st ri on ics L abo ra to ry

C om pu ta ti on L abo ra to ry

A st ro na ut ics L abor at or y

P ro gr am Ma na ge me nt

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 Rockwell/Rocketdyne Division

xxiii



A B B R E V I A T I O N S

ACN Ascension DTO Detailed Test Objective

AEDC Arnold Engineering EBW Exploding Bridge Wire

Developmententer

E C O E n g in e C u to ffA G C A utom a tic G a in C o ntro l

E C P E ng ineering C hange P roposa lANT Antigua

E C S E nvironm enta l C ontro l S ystemA O S A cqu isition o f S igna l

E D S E m ergency D etection S ystemA P S A uxilia ry P ropu ls ion S ystem

E D T E astern Dayligh t T im eA S l A ugm en ted S park Ign ite r

E M R E ng ine M ixtu re R a tio

A U X A u xilia ryE S C E ng ine S ta rt C om m an d

BDA Bermuda

E VA E xtra -Vehicu la r A ctivityC C S C om m and and C om m unica tions

System FCC Flight Control Computer

CDDT Countdown Demonstration Test FM/FM Frequency Modulation/F requency M odulation

C E C O C ente r E n g ine C u to ffG B I G ra nd B ah am a Is la n d

CG Center of Gravity

G B M G ra nd B ah am a Is la n d

C KA F S C ape Kennedy A ir F orce S iteG F C V G O X F low C on tro l V a lve

CM CommandModule

G D S G o ldsto n eC N V C a navera l

GG Gas GeneratorC R O C a rna rv o n

G M T G re en w ich M e an T im eC R P C o m pu ter R eset P u lse

G O X G a se o us O xyg enC S M C om m and a nd S erv ice M odu le

G R R G uidan ce R efe ren ce R elea se

C VS C on tinuous V en t S ystemG S E G rou nd S u ppo rt E qu ipm en t

C Y I G rand C anary Is la nd

G T I G ra n d T u rk Is la n dD E E D ig ita l E ven ts E va lua to r

GWM Guam

xxiv



GYM Guaymas MER Mercury(ship)

HAW Hawaii MFCV Modulating Flow Control Valve

HDA HolddownArm MFO Major Flight Objective

HEP Hardware Evaluation Program MFV Main Fuel Valve

HFCV Helium Flow Control Valve MILA Merritt Island Launch Area

HSK Honeysuckle (Canberra) MLV Main LOX Valve

IGM Iterative Guidance Mode MOV Main Oxidizer Valve

IMU Inertial Measurement Unit MR Mixture Ratio

IP & C In stru m e n ta tio n P ro g ram an d M S C M a nn ed S p acecra ft C e n terComponents

M S F C M arsha ll S pace F ligh t C enterIU Instrument Unit

M S F N M anned S pace F ligh t N etw orkK S C K en n edy S p a ce C en te r

M S S M o bile S erv ice S truc tu reL E S L a u n ch E sca p e S ystem

M T F M iss issipp i T est F ac ilityLET Launch Escape Tower

N P S P N et P ositive S uction P ressureL IE F L aunch In fo rm ation E xchange

Facility NPV Non Propulsive Vent

LM Lunar Module NASA National Aeronautics andS p ac e A d min is tra tio n

L O I L u na r O rb it In sertio n

OAFPL Overall Fluctuating PressureLOS Lossof Sgnal Level

LUT Launch Umbilical Tower OASPL Overall Sound Pressure Level

LV Launch Vehicle OAT Overall Test

L V D A L au nch V eh ic le D a ta A da p te r O C P O rb ita l C o rre ction P rog ra m

LVDC Launch Vehlcle Digital ODOP Offset Frequency Doppler

Computer

O E C O O u tboard E n g ine C u to ffMAD Madrid

O M N I O mni D irectiona lM C C -H M iss io n C on tro l C en te r -

Houston

XX V



PAM/ Pulse Amplitude Modulation/ SPL Sound Pressure LevelFM/FM Frequency Modulation/

Frequency Modulation SRSCS Secure Range Safety CommandSystem

PAFB Patrick Air Force Base

SS/FM Single Sideband/FrequencyPCM Pulse CodeModulation Modulation

PCM/ Pulse Code Modulation/ STDV Start Tank Discharge ValveF M F re quency M odu la tion

SV Space Vehicle

P M R P rog ram ed M ixture R a tio

T1 Time Base 1P R A P atrick R eference A tm osphere

TII Time to go in Ist Stage IGMP S D P ow er S pectra l Density

T21 Time to go in 2nd Stage IGMP T C S P rop ellan t T ank ing C ontro l

System TAN Tananarive

PU Propellant Utilization TCS Thermal Conditioning System

RED Redstone (ship) TD&E Transposition, Docking andEjection

RF Radio Frequency

T E L 4 C ape T e lem e try 4R M S R o o t M ea n S qu a re

TEX Corpus Christi (Texas)R P -I D esign a tio n fo r S -lC S tag e

Fuel (kerosene) TLI Translunar Injection

RPM Revolutions Per Minute TM Telemeter, Telemetry

RTG Radio Isotope Thermo- TMR Triple Modular RedundantE lectrical Generator

T S M T a il S e rv ice M a stSA Service Arm

T VC T hrust Vecto r C on tro l

SC Spacecraft

U H F U ltra H igh F reque ncySEC Seconds

U S B U n ifie d S -B a nd

S L A S pacecra ft L M A dap te r

UT Universal TimeSM Service Module

VAN Vanguard (ship)S M C S teering M isa lignm ent

Correction VHF Very High Frequency

WHS White Sands

xxvi



M IS S IO N P L A N

T he A S -50 5 (A po llo I0 m ission) is the fifth fligh t o f the A po llo-S aturn Vflight test program . It is a L unar Developm ent F light ; the prim aryobjectives are: (I) demonstrate crew/space vehicle/mission supportfacilities performance during a manned lunar m ission with the Command andS ervice M odule (C S M ) and L unar M odule (L M ); (2) evaluate L M perform ancein the cislunar and lunar environm ent. T he crew is com posed ofL t. Col. T homas S tafford, Cdmr. John Y oung and Cdmr. E ugene Cernan.

T he space vehicle is com posed of the A S -50 5 L aunch Vehicle (L V ) consisting

of the S -IC-5, S -II-5, S -IV B -5 and Instrum ent Unit (IU) stages andspacecraft consisting of the S pacecraft L M A dapter (S L A ), L M -4 and CS M -I0 6 .

The vehicle is launched from Complex 39B at Kennedy Space Center.T he launch azim uth is 9 0 degrees with a ro ll to a variable flight azim uth

of 7 2 to 1 0 8 degrees east o f true north .

The vehiclemass at launch(GroundIgnition)is about2,945,069kilograms(6,492,766 Ibm). The S-IC and S-II stage powered flight times are approx-

imately 16O and 392 seconds, respectively. The S-IVB first burn time is

approximately145 seconds. The S-IVB/IU/LM/CSMisinsertedinto a185 kilometer (lO 0 n mi) altitude (referenced to the earth's equatorial

radius) circular parking orbit. The vehicle mass at parking orbit

i nser tio n is abo ut 13 3,760 ki logr ams ( 294,891 Ibm) .

A bout lO seconds after insertion into earth orbit, the vehicle assumes

a horizontal attitude. During this coast in earth orbit, the LV and CSM

system is checked out for T ranslunar Injection (TLI). During the

second or third revolution the second burn (344 seconds) of the S-IVB

injectsthe S-IVB/IU/LM/CSMinto a free-return,translunartrajectory.

Fifteen minutes after S-IVB cutoff, the LV maneuvers to an inertial

attitude hold for CSM separation, docking and LM extraction. After the

maneuver, the CSM separates from the LV and the SLA panelm jettison.

The CSM then transposesand docks to the LM. After docking,the CSM/LM

is ejected,by springs,fromthe S-IVB/IU.

After the CSM/LMhas been ejected,the S-IVB stage achievesa slingshot

trajectory behind the moon and a solar orbit by activating propulsion

venting, the contingency experiment of propellant lead, LOX dump through

the J-2 engine, and firing the Auxiliary Propulsion System (APS) ullage

engines.

x x v i i



During the 3-day translunar coast of the CSM/LM, the astronauts performstar/earth landmark sighting, Inertial Measurement Unit (IMU) alignmentsand general lunar navigation procedures. A t approxim ately 7 6 .5 hours,

a Lunar Orbit Insertion (LOI) burn puts the CSM/LM into a III by 315k ilom eter (6 0 by 1 7 0 n m i) e llip tica l o rbit by a n a pprox im ate 380 -seco nd

S erv ice P ropu ls ion S ystem (S P S ) burn . A fter tw o revo lu tions, the C S M

circu la rizes the o rbit a t I ll k ilom eters (6 0 n m i) by a 1 5-second S P S

burn . T he L M is then entered by two astronauts and checkou t accom plished.

A t approxim ately 9 8.8 hours, C S M undocking occurs. A t I0 1 hours, a

Descent P ropu ls ion S ystem (D P S ) phasing burn p laces the L M in a 39 4 by

18 kilometer (213 by 9.9 n mi) orbit. The LM simulates the descent atti-tude profile during approach to the phasing burn. A t approxim ately 1 0 3

hours, an A P S burn in itia tes L M active rendezvous; L M docking occurs a t

approxim ately 1 0 7 hours fo llowed by L M deactiva tio n and crew transfer to

th e C S M . T he L M is je ttison ed a t app ro xim ate ly 1 0 9 ho urs a nd the C S M in -

jected in to the transearth tra jectory. T h e coast period lasts approxi-

m a te ly 89 hou rs . T he S erv ice M odu le (S M ) separa tes fro m the C o m m an dM odu le (C M ) 1 5 m in u tes prio r to reen try . S p la shdow n in the P ac ific O cean

occurs approximately 1 91 hours after lifto ff.

xxviii



FL IGHT T E S T S UM M A R Y

T he third manned S aturn V A pollo space vehicle, A S -50 5 (A pollo I0 M ission),was launched at Kennedy S pace Center (KS C), F lorida on M ay 1 8, 1 9 6 9 at

1 2:49 :0 0 E astern Daylight T im e (E DT ) from L aunch Com plex 39 , P ad B . T hisfifth launch of the Saturn V Apollo was the second Saturn V/ApolloS pacecraft in "fu ll lunar mission configuration. T he three major flightobjectives and the six Detailed T est O bjectives (DT O 's) were com pletelyaccomplished.

The launch countdown was completed without any unscheduled countdown

holds. Ground system performance was satisfactory. T he problemsencountered during countdown were overcome such that vehicle launchreadiness was not compromised.

T he vehicle was launched on an azim uth of 9 0 degrees east of north andafter 1 3.0 5 seconds of vertical flight, the vehicle began to roll in to

a flight azim uth of 72.0 28 degrees east of north . A ctual tra jectoryparameters of the A S -50 5 were close to nominal. S pace-fixed velocityat S-IC Outboard Engine Cutoff (OECO)was 10.81 m/s (35.47 ft/s) greaterthan nominal. At S-II OECOthe space-fixed velocity was 13.22 m/s(43.37 ft/s) lower than nominal. At S-IVB first cutoff the space-fixedvelocitywas 0.07 m/s (0.23 ft/s) greater than nominal. The altitude atS -IVB first burn cuto ff was 0 .0 3 kilom eters (0 .0 1 n m i) lower than

nom inal, and the surface range was 0 .9 2 kilom eters (0 .50 n m i) greaterthan nom inal. T he space-fixed velocity at parking orbit insertion was0.07 m/s (0.23 ft/_less than nominal. At translunar injection thetotal space-fixed velocity was 2.39 m/s (7.84 ft/s) less than nominal.

The value of C3 was 868 m_/s2 (9345 ft2/s 2) lower than nominal.

A ll S -IC propulsion systems performed satisfactorily. A t the 35 to38-second time slice, average engine thrust reduced to standard conditionswas 0 .20 percent lower than predicted. A verage reduced specific impulsewas 0.03 percent lower than predicted, and reduced propellant consumptionra te was 0 .1 58 percent lower than predicted. Center E ng ine C utoff (C E CO )was in itia ted by the Instrum ent Unit (IU) at 1 35.1 6 seconds as planned.

O E CO , initiated by L O X low level sensors, occurred at 1 61 .63 secondswhich was 1 .43 seconds later than predicted in the F light T ra jectory.

The S -II propulsion system performed satisfactorily throughout the flight.A s sensed by the engines, E ngine S tart Com m and (E S C) occurred at 1 6 3.0 5seconds. O E CO occurred at 552.6 4 seconds with a burn tim e of 389 .59

seconds or 1 .70 seconds longer than predicted. Due to center engine low

xxix





The structural loads and dynamic environments experienced by the A S -505

launch vehicle were well with in the vehicle structural capability. Therewas no evidence of coupled structure/propulsion system instability (POGO)during S-IC, S-II, or S-IVB powered flights. The early S-II stage centerengine shutdown successfu lly eliminated the low frequency (1 6 to 1 9 hertz)oscillations that were experienced on A S -50 3 and A S -50 4. During S -IVB

first and second burns, very m ild low frequency (1 2 to 1 9 hertz) oscilla-tions were experienced with the maximum amplitude of ±0 .30 g recorded bythe gimbal block longitudinal accelerometer. During the last 7 0 secondsof second burn, the A pollo I0 astronauts reported (in rea l tim e) that

higher frequency oscillations were superimposed on the low frequencyoscilla tions. These vibrations are, however, well with in the structuraldesign capability.

The guidance and navigation system functioned satisfactorily. T ranslunar

trajectory injection parameters were within tolerance, and S -IVB stagesafing was satisfactorily accomplished, resulting in a heliocentric orbit

for the S-IVB/IUas planned. The LVDC, the Launch Vehicle Data Adapter

(LVDA ), and the S T-1 24M -3 inertia l platform functioned satisfactorily.

T he A S -50 5 F light Contro l Com puter (F CC), T hrust Vector Contro l (T VC), and

A P S satisfied all requirements for vehicle attitude contro l during theflight. S-IC/S-II first and second plane separations were accomplishedwith no significant attitude deviations. A t S -II planned CE CO , theguidance parameters were modified by the loss in thrust. S-II/S-IVBseparation occurred as expected and without producing any significantattitude deviations. S atisfactory contro l of the vehicle was maintained

during first and second S -IVB burns and during parking orbit. During theCommandand Service Module (CSM) separation from the S-IVB/IU and duringthe T ransposition , Docking and E jection (T D&E ) maneuver, the control systemmaintained'the vehicle in a fixed inertial attitude to provide a stable

docking pla tform. A fter T ranslunar Injection (T L I), attitude control wasmainta ined for the propellant dumps and chilldown experim ent. F or A S -50 5the A P S propellants were not depleted by the last u llage burn, and controlwas mainta ined until the batteries were exhausted.

The A S -505 launch vehicle electrical systems performed satisfactorilythroughout all phases of flight. Data indicated that the redundantS ecure R ange S afety Com m and S ystem s (S R S CS ) on the S -lC , S -II , and S -IVBstages were ready to perform their functions properly on command if flightconditions during the launch phase had required vehicle destruct. T hesystem properly safed the S -IVB S R S CS o n com m and from B erm uda (B DA ). T he

perform ance of the C om m and and Com m unications S ystem (CC S ) in the IU was

satisfactory, except during the tim e period from 23,60 1 seconds (0 6 :33:21 )when CCS downlink signal strength dropped sharply until 25,0 9 7 seconds(0 6 :58:1 7 ), when the antenna was switched to the om ni m ode. T he drop insignal strength is suspected to be a malfunction in the directionalantenna system.

xxxi



T he E mergency Detection S ystem (E DS ) perform ance was nominal; no abort

lim its were reached. T he A S -50 5 E DS configuration was essentially thesame as A S-504.

The vehicle internal, external, and base region pressure environments

were generally in good agreement with the predictions and compared wellwith previous flight data. T he pressure environm ent was well belowdesign levels. T he measured acoustic levels were genera lly in goodagreement with the liftoff and inflight predictions, and with data from

previous flights.

T he A S -50 5 vehicle thermal environment was similar to that experienced

on earlier flights with the exception of the S -IC stage which showedminor changes due to differences of higher ambient temperatures atl i f t o f f .

T he E nvironmental Contro l S ystems (E CS ) performed satisfactorily duringthe A S -50 5 countdown. A vailable data shows the IU E CS performed

satisfactorily. The IU environmental conditioning purge duct exhibited apressure loss and flow increase during prelaunch operations but IUperformance was unaffected.

All elements of the data system performed satisfactorily except for aproblem with the CCS downlink during translunar coast. M easurementperformance was excellent, as evidenced by 99.2 percent reliability.T his is the highest reliability atta ined on any S aturn V flight.T elemetry performance was nominal, with the exception of a minorcalibration deviation. T he onboard tape recorder performance wassatisfactory. V ery H igh F requency (V HF ) telem etry R adio F requency (R F )propagation was generally good; though the usual problems due to flameeffects and stag ing were experienced. VHF data were received to1 5,7 80 seconds (0 4:23:0 0 ). Comm and system s R F perform ance for boththe S R S CS and CCS was nominal except for the CCS downlink problemnoted. G oldstone (G DS ) and G uaym as (G Y M ) reported receiving CCSsignal to 40 ,1 91 seconds (1 1 :0 9 :51 ). Good tracking data were receivedfrom the C-B and radar with B DA indicating fina l L oss of S ignal (L O S )at 35,346 seconds (0 9 :49 :0 6 ). T he 7 3 ground engineering cam erasprovided good data during the launch.

x x x i i



S E CT IO N l

INTRODUCTION

l.l P U R P O S E

T his report provides the N ationa l A eronautics and S pace A dmin istration

(N A S A ) H eadquarters, and o ther interested agencies, w ith the launch vehi-

c le eva lua tion resu lts o f the A S -50 5 fligh t test. T he basic objective o f

fligh t eva lu ation is to acqu ire, reduce, ana lyze, eva luate and rep ort on

flight test data to the extent requ ired to assure fu tu re m ission success

and veh icle reliability. T o acco m plish th is o bjective, a ctua l fligh tfa ilures, anom alies and devia tions m ust be identified, their causes ac-

cura te ly determ ined, and com plete in formation m ade ava ilable so that

corrective action can be accomplished with in the established flightschedule.

] .2 S C O P E

T h is report presents the resu lts o f th e early eng ineering flig h t eva lua -

tion o f the A S -50 5 launch veh icle . T he con ten ts are cen tered on the per-

fo rm ance evaluation of the major launch veh icle systems, with specia lem phasis on fa ilu res, anom alies, and dev ia tion s. S um m aries of launch

operations and spacecraft performance are included for completeness.

T he offic ia l G eorge C . M arshall S pace F ligh t C enter (M S F C ) position a t

th is tim e is represented by th is repo rt. I t w ill no t be fo llow ed by asim ilar report un less continued ana lysis or new information shou ld prove

the conclusions presented herein to be sign ificantly incorrect. F ina l

sta ge eva luation reports w ill, however, be p ublished by th e stage con-tracto rs. R epo rts covering m ajor subjects a nd specia l subjects will be

published as required.

I-I/I-2



S E CT IO N 2

E VE NT T IM E S

2.1 S U M M A R YO F E V E N T S

R ange zero tim e, the basic tim e reference fo r th is report, is 1 2:49 :0 0E astern D ayligh t T im e (E D T ) (1 6 :49 :0 0 U n iversa l T im e [U T ]). T h is tim e

is based on the nearest second prio r to S - IC ta il p lug disconnect wh ich

occurred a t 1 2::49 :0 0 .6 E DT . R ange tim e is ca lculated as the elapsedtim e from range zero tim e and, un less o therwise no ted, is the tim e used

throughout this report. The actual and predicted range times are ad-

justed to ground te lem etry received tim es. F igure 2-I shows the tim ede lay of g ro und telem etry received tim e versus L aunch V ehic le D ig ita l

C om puter (L V DC ) tim e and indica tes the m agn itude and sign o f correctionsapp lied to correla te range tim e and veh icle tim e in T ables 2- I, 2-2 and2-3.

G uidance R eference R elease (G R R ) occurred a t -1 6 .9 7 seconds and sta rt o f

T im e B ase 1 (T I) occu rred a t 0 .58 secon ds. G R R w as es tablish ed by the

D ig ita l E ven ts E va lua to r (D E E -6 ) and T 1 was in itia ted a t detection o f

lifto ff signa l prov ided by de-energ izing the lifto ff re lay in the In stru -m en t U n it (IU ) a t IU um bilica l d isco nnect.

R a ng e tim e fo r ea ch tim e ba se used in th e fligh t sequ ence program an d the

signa l for in itia ting each tim e base are presen ted in T able 2-I.

S tart o f T 2 was with in nom ina l expecta tions fo r th is even t. S ta rt o f T 3,T 4 and T 5 w as in itia ted approx im ate ly 1 .5 , -1 .4 and 0 .3 seconds later than

predicted, respectively, due to variations in the stage burn tim es. T hesevaria tio ns are discussed in S ections 5, 6 , and 7 of th is docum en t. S ta rt

o f T 6 , w h ich w as in itia ted by the L VD C upon so lv ing the resta rt equa tion ,

w as 2.4 seco nds la ter tha n predicted. S ta rt o f T 7 w a s 2.0 seconds la tertha n pred ic ted. T 8, w h ich w as in itia ted by the receip t o f a g ro und com -

m a nd, was started 1 86 .8 seco nds la ter than the predicted tim e.

A su m m ary o f s ig n ifican t even ts fo r A S -50 5 is g iven in T a ble 2-2 . S inceno t a ll even ts listed in T able 2-2 are IU com m anded switch selecto r func-

tions, deviations are not to be construed as failures to meet specifiedswitch selecto r to lerances. T he events in T able 2-2 associa ted with

gu idance, navigation, and control have been identified as being accura teto with in a m ajor com puta tion cycle.

2 - I



320

_S-IVB SECOND BURN /FU EL LEA D EX PER IM EN T

280 /

0U

240

E

200

W

160 /

120

Z

_ 80

40

O.0 4 8 12 16 20 24 28

RANGE T IM E, 1000 SECONDS

i i i _ i i _7, t i0:00:00 1:00:00 2:00:00 3:00:00 4:00:00 5:00:00 6:00:00 7:00:00

RANG E T IM E , HOU R S:M IN U TE S:SE COND S

Figure 2-I. Telemetry Time Delay

2-2



T able 2- I. T im e B ase S um m ary

P _ A N G E T I M E

S EC

TIMEBASE (HR:MIN:SEC) SIGNALSTART

TO -16.97 Guidance Reference Release

T1 0.58 IU Umbilical Disconnect SensedbyLVDC

T2 135.29 S-IC CECOSensedby LVDC

T3 161.66 S-IC OECOSensedby LVDC

T4 552.65 S-II ECOSensedby LVDC

T5 703.98 S-IVB ECO(Velocity) SensedbyLVDC

T6 8629.26 Restart Equation Solution(2:23:49.26)

T7 9550.83 S-IVB ECOCommandedy LVDC(2:39:10.83)

,935.83 Enabled by Ground CommandT 8 (4: :IS 83)

T he predicted tim es fo r establish ing actua l m inu s predicted tim es inT able 2-2 have been ta ken from 40 M 336 25, " In terface C ontro l Docum ent

Definition of Saturn SA-505 Flight Sequence Program" and from the "SaturnV A S -50 5 P ost-L aunch P redicted O perationa l T ra jectory," dated M ay 23, 1 96 9 .

2.2 VA R IA B LE T IM E A N D C O M M AN DE DS W IT C H S E L E C T O R E VE N TS

T able 2-3 lists known switch selector events which were issued during

flight but which were not programed for specific times. The water coolant

va lve open and close switch selector com m ands were issued based upon thecondition o f two therm al switches in the E nvironm enta l C ontro l S ystem (E C S ).

T he ou tputs o f these switches were sam pled once every 30 0 seconds, begin-

n ing a t 1 80 seco nds, a nd a sw itch selecto r com m a nd w a s issu ed to o pen an dclose the water va lve to m ain ta in proper tem perature con tro l.

2-3



Table 2-2. S ignificant E vent T imes S ummary

RANGE TIME TIME FROM BASE

EVENT DESCRIPTION ACTUAL ACT-PRED ACTUAL ACT-PRED

SEC SEC SEC SEC

I GUIDANCE REFERENCE RELEASE -17.0 0.0 -17.6 0o0

(GRR)

2 S-IC ENGINE START SEQUENCE -8.9 0.0 -qo5 0;ICOMMAND (GROUND)

3 S-IC ENGINE N0.1 START -6.1 0.0 -6.7 0,0

4 S-IC ENGINE NO.2 START -6.0 0.0 -6.6 0.0

5 S-IC ENGINE N0.3 START -6,3 0.1 -6.9 0.1

6 S-IC ENGINE N0.4 START -6.0 0.0 -6.6 0.0

7 S-IC ENGINE N0.5 START -6.4 0.0 -7.0 0,0

8 ALL S-IC ENGINES THRUST OK -1.6 -0.I -2.2 -O°l

9 RANGE ZERO 0.0 -0.6

10 ALL HOLDDOWN ARMS RELEASED 0.2 0.0 -0.3 0.I

(FIRST MOTION)

II IU UMBILICAL DISCONNECT_ START 0.6 0o0 0.0 0,0

OF TIME BASE I (TIT

12 BEGIN TOWER CLEARANCE YAW 1.6 0.0 1.0 0.0

MANEUVER

13 END YAW MANEUVER IO.O 0.4 9.4 0.4

14 BEGIN PITCH AND ROLL MANEUVER 13,1 0.6 12.5 0.6

15 S-IC OUTBOARD ENGINE CANT 20.6 0°0 20.0 O.O

16 END ROLL MANEUVER 32.3 1.8 31.7 1.8

17 MACH I 66.8 0,9 66.2 I.0

IB MAXIMUM DYNAMIC PRESSURE 82.6 1.5 82,0 1.5

(MAX Q)

19 S-IC CENTER ENGINE CUTOFF 135.2 -0.I 134.6 O.O

(CECO)

20 START OF TIME BASE 2 (T2) 135.3 0.I 0._ 0.0

21 END PITCH MANEUVER ITILT 158.7 2.0 23.4 1,9

ARREST)

22 S-[C OUTBOARD ENGINE CUTOFF 161.6 I.4 26.3 1.4

(OECO|

23 START OF TIME BASE 3 |T3) 161°7 I.5 0.0 0.0

24 START S-11 LH2 TANK HIGH 161.7 1.4 0. l 0.0PRESSURE VENT MODE

2 - 4



T able 2-2. S ignificant E vent T im es S umm ary (Continued)


EVENT DESCRIPTION ACTUAL ACT-PREU ACTUAL ACI-PREO

SEC SEC SEC SEC

25 S-If LH2 RECIRCULATION PUMPS 161.8 1.4 0.2 0.0OFF

26 S-IT ULLAGE MOTOR IGNITION 162.1 1.4 0.5 0.0

27 S-ICIS-II SEPARATION COMMAND 162.3 1.4 0.7 O.O

TO FIRE SEPARATION DEVICES

AND RETRO MOTORS

28 S-IC RETRO MOTOR EFFECTIVE 162.4 1.4 0.8 0.0BURN TIME INITIATION (THRUST

BUILDUP REACHES 75%)

(AVERAGE OF 8)

Z9 S-IT ENGINE START COMMAND 163.1 1.5 1.4 O.O

IESC)

30 S-II ENGINE IGNITION ISIDV 164.1 1.5 2.4 O.OOPEN, AVERAGE OF FIVE)

31 S-IT ULLAGE MOTOR BURN TIME 166.0 1.2 4.4 -0.2

TERMINATION (THRUST REACHES

75_)

32 S-IT MAINSTAGE 166o3 1.7 4.6 0.2

33 S-If CHILLDOWN VALVES CLOSE 168.0 1.4 6.4 O.O

34 ACTIVATE S-IT PU SYSTEM 168.5 [.4 6.9 0.0

35 S-If SECOND PLANE SEPARATION [92.3 1.4 30.7 O.OCOMMAND (JETTISON S-IT AFT

INTERSTAGEI

36 LAUNCH ESCAPE TOWER (LET) 197.8 1.4 36.1 -0.9JEITISON

37 ITERATIVE GUIDANCE MODE IIGM) 202°9 1.4 41.2 -O.l

PHASE [ INITIATED

3B S-II LOX STEP PRESSURIZATION 261.6 1.4 [OO.O 0.0

39 S-II CENTER ENGINE CUTOFF 460.6 1.4 299.0 0.0

(CECO)

40 S-If LH2 STEP PRESSURIZATION 461.6 1.4 300.0 0o0

4I GUIDANCE SENSED TIME TO BEGIN 484.8 1.5 323°I O.I

EMR SHIFT IIGM PHASE 2 INI-

TIATED & START OF ARTIFI-

CIAL TAU MODE)

42 S-If LOW ENGINE MIXTURE RATIO 488.5 0.0 326.8 -[.5(EMR) SHIFT IACTUAL)

43 END OF ARTIFICIAL TAU MODE 490.2 6.7 328.6 5.3

2 - 5



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

RANGE TIME TIME FROM BASEEVENT DESCRIPTION ACTUAL ACT-PRED ACTUAL ACT-PRED

SEC SEC SEC SEC44 S-II OUTBOARD ENGINE CUTOFF 552.6 -1.5 391.0 -2.9

(OECU)

4§ S-IT ENGINE CUTOFF INTERRUPT, 552.7 -1.4 0.0 0.0

START OF TIME BASE 4 (r4)

(START OF IGM PHASE 3)

46 S-IVB ULLAGE MOTOR IGNITION 553.4 -1.5 0.8 0.0

47 S-II/S-IVB SEPARATION COMMAND 553.5 -I.5 0.9 0.0

TO FIRE SEPARATION DEVICES

AND RETRO MOTORS

4B S-IVB ENGINE START COMMAND 553.6 -1.5 1.0 O.O

(FIRST ESC)

49 FUEL CHILLDOWN PUMP OFF 554.8 -1,5 2.2 0.0

50 S-IVB IGNITION ISTDV OPEN) 556.g -|.Z 4.2 0.2

51 S-IVB MAINSTAGE 559°3 -1.3 6.7 0.2

52 START OF ARTIFICIAL TAU MODE 560.I -1.9 7.5 -0.4

53 S-IVB ULLAGE CASE JETTISON 565.4 -1.5 12.8 0.0

54 END OF ARTIFICIAL TAU MODE 568.9 -3.6 16.2 -2.2

55 BEGIN CHI BAR STEERING 669.4 0.6 116.7 2.1

56 END IGM PHASE 3 695.7 -0,6 143.1 1.0

57 BEGIN CHI FREEZE 697.3 1.0 144.6 2.5

58 S-IVB VELOCITY CUTOFF COMMAND 703.8 0.3 151.[ 1.8

(FIRST GUIDANCE CUTOFF}

(FIRST ECO)

59 S-IVB ENGINE CUTOFF INTERRUPT. 704.0 0.3 O.O O.O

START OF TIME BASE 5 (75)

bO S-IVB APS ULLAGE ENGINE NO. I 704.3 0.3 0.3 0.0IGNITION COMMAND

61 S-IVB APS ULLAGE ENGINE NO. 2 704.4 0.3 0.4 0.0

IGNITION COMMAND

62 LOX TANK PRESSURIZATION OFF 705.3 0.2 [.4 0.0

63 PARKING ORBIT INSERTION 7t3.B 0.3 9.8 0.0

64 BEGIN ORBITAL GUIDANCE, BEGIN 724.[ 0.2 20. I -O.I

MANEUVER TO LOCAL HORIZONTAL

ATTITUDE

65 S-IVB LH2 CONTINUOUS VENT 763.0 0.3 59.0 0.0

SYSTEM ICVS} ON

2 - 6



T able 2-2. S ign ificant E vent T im es S um m ary (Continued)


EVENT DESCRIPTION ACTUAL ACT-PRED ACTUAL ACT-PRED!

SEC SEC SEC SEC

66 S-IVB APS ULLAGE ENGINE NO. l 791.0 0.3 87.0 O.OCUTOFF COMMAND

67 S-IVB APS ULLAGE ENGINE NO. 2 791.0 0.2 87.1 0.0CUTOFF COMMANO

68 FIRST ORBITAL NAVIGATION 805.2 -6.7 101.2 -6.8CALCULATIONS

69 BEGIN S-IVB RESTART PREPARA- 8629.] 2.4 O.O O.O

TIONS, START OF TIME BASE 6

(T6)

70 S-IVB 02/H2 BURNER LH2 ON 8670.5 2.3 41.3 O.O

71 S-IVB 021H2 BURNER EXCITERS ON 8670.8 2.3 41.6 O.O

72 S-IVB D21H2 BURNER LOX ON 8671.2 2.3 42.0 0.0{HELIUM HEATER ON)

73 S-IVB LH2 VENT OFF (CVS OFF) 8671.4 2.3 42.2 0o0

74 S-IVB LH2 REPRESSURIZATION 8677.3 2.3 48.I 0.0

CONTROL ON

T5 S-IVB LOX REPRESSURIZATION 8677.5 2.3 48.3 0.0

CONTROL ON

76 S-IVB AUX HYDRAULIC PUMP 8848.2 2.3 219.0 0o0FLIGHT MODE ON

77 S-IVB LOX CHILLDOWN ON 8878.2 2.3 24g°0 O.O

78 S-IVB LH2 CHILLDOWN ON 8883.2 2.3 254.0 0.0

79 S-IVB PREVALVES CLOSED 8888.2 2.3 259.0 O.O

80 S-IVB PU MIXTURE RATIO 4.5 ON 9079.3 2°3 450.I 0.0

81 S-IVB APS ULLAGE ENGINE NO. I 9125.5 2.3 496.3 0.0IGNITION COMMAND

82 S-IVB APS ULLAGE ENGINE NO. 2 9]25.6 2.3 496.4 O.OIGNITION COMMAND

83 S-IVB 02/H2 BURNER LHZ OFF 9126.0 2.3 496°8 0.0(HELIUM HEATER OFF)

84 S-IVB 021H2 BURNER LCX OFF g130.5 2.3 501.3 0.0

85 S-IVB LH2 CHILLDOWN OFF 9198.6 2.3 56g.4 0.0

86 S-IVB LOX CHILLDOWN OFF 9198.B 2,3 56g°_ 0.0

87 S-IVB ENGINE RESTART COMMAND 9199.2 2.3 510.0 O.O(FUEL LEAD INITIAIIGN)

(SECUND ESC)

2 - 7



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


EVENT DESCRIPTION ACTUAL ACT-PRED ACTUAL ACT-PREDSEC SEC SEC SEC

88 S-[VB APS ULLAGE ENGINE NO. I 9202°2 2.3 573.0 0.0

CUTOFF COMMAND

89 S-[VB APS ULLAGE ENGINE NO. 2 9202.3 2°3 573.1 0.0

CUTOFF COMMAND

90 S-IVB SECOND IGNITION (STDV 9207.5 2.6 578.3 0.3

OPEN)

91 S-IVB MAINSTAGE 9210.0 2.6 580.8 2.3

92 IGM PHASE 4 INITIATED 9218.2 6.8 589.0 4.5

93 ENGINE MIXTURE RATIO (EMR) 9334.3 I.O 705.I -1.3

SHIFT

94 S-IVB LH2 STEP PRESSURIZATION 9479.2 2.3 850.0 0.0

(SECOND BURN RELAY OFF)

95 BEGIN CHI BAR STEERING 9521.1 l.T 891°9 -0.6

96 BEGIN CHI FREEZE 9549.3 2.4 920.1 0o1

9T S-[VB SECOND GUIDANCE CUTOFF 9550.6 2°0 -0.3 -O°l

COMMAND (SECOND ECO)

98 S-IVB ENGINE CUTOFF INTERRUPT, 9550.8 2.0 0.0 0.0

START OF TIME BASE 7

99 LH2 VENT ON COMMAND 9551.3 2.0 0.5 0.0

100 TRANSLUNAR INJECTION 9560.6 2.0 9.7 -0.I

lOT BEGIN ORBITAL GUIDANCE 9569.6 0.I 18.9 -I.8

102 FIRST ORBITAL NAVIGATION 9572°5 2.0 21.7 0.0

CALCULATIONS

103 CSM SEPARATION 10962.4 -42.5 1411.5 -44.6

104 CSM DOCK 11856.0 416.0 2305. I 414,0

105 SCILV FINAL SEPARATION 14185.7 -719.3 4634.8 -721.3

106 INITIATE MANEUVER TO SLINGSHOT 16935.8 186o8 0.0 0.0

ATTITUDE

I07 START OF TIME BASE 8 {TB} 16935.8 IBb.8 0.0 0.0

108 S-IVB ULLAGE ENGINE NO. I ON 17136.4 187.4 200.6 0°6

109 S-IVB ULLAGE ENGINE NO. 2 ON I7137.B I88.B 202.0 2.0

IlO BEGIN LOX REPRESSURIZATION 17153.I 189oi 217.3 2.3

1[1 BEGIN LOX LEAD 17301.3 187.3 365.5 0.5

2 - 8



T ab]e 2-2. Significant E vent T imes Summary (C ontinued)

RANGE TIME TIME FROM BASEEVENT DESCRIPTION ACTUAL ACT-PRED ACTUAL ACT-PREO

SEC SEC SEC SEC

112 END LOX LEAD _7310.3 188.3 376°5 [,5

_13 END COX REPRESSURIZATION 17355.5 268.5 419,6 81,6

11¢ BEGIN LH2 REPRESSURIZATIDN 17356.9 205.9 421.0 19o0

ll5 END LH2 REPRESSURIZATIDN 17386.2 175.2 450.4 -11.6

116 BEGIN FUEl. LEAD 17409.9 187.9 474.1 l.l

117 S-IVB ULLAGE ENGINE NO. I OFF 17415.6 190.6 479.7 3°7

118 S-IVB ULLAGE ENGINE NO. 2 OFF 17417.0 192.0 481.1 5.1

119 END FUEL LEAD [7458.B 183.8 523.0 -3,0

[20 H2 CONTINUUUS VENT ON 17496.1 206.1 560.2 I7.2

121 BEGIN LOX DUMP 17655.8 186.8 720.0 O°O

122 END LOX DUMP 17956.0 186o8 1020.2 0°0

123 H2 NONPROPULSIVE VENT ON (NPV] 18969.8 186,8 2034.0 0o0

124 S-IVB ULLAGE ENGINE NO. I ON I9735.8 186.8 2800.0 O,O

I25 S-IVB ULLAGE ENGINE NO, 2 ON 19736.0 I87o0 2800,2 0°2

126 S-IVB ULLAGE ENGINE NO. I OFF 19743.3 39.3 2807.4 -147.6

127 S-IVB ULLAGE ENGINE NO. 2 OFF 19744.9 40.9 2809.0 -146.0

2 - 9



T able 2-3. Variable T ime and Commanded S witch S elector E vents

RANGETIME TIME FROMBASE

FUNCTION STAGE (SEC) (SEC) REMARKS

Water Coolant Valve Open IU 180.8 T3 +19.2 LVDCFunction

Water Coolant Valve Closed IU 782.5 T5 +78.5 LVDC Function

Start Calibration Sequence 2293.3 T5 +1589.4 TANRev l

Start Calibration Sequence 3205.3 T5 +2501.3 CRORev1

Start Calibration Sequence 5373.4 T5 +4669.4 GYMRev1

Start Calibration Sequence 6677.4 T5 +5973.4 CYI Rev 1

Start Calibration Sequence 7821.4 T5 +7117.3 TANRev2

Water Coolant Valve Closed IU 8000.0 T5 +7296.0 LVDCFunction

Start Calibration Sequence 9029.2 T6 +400.0 CRORev2

Ambient Repress ModeSelector Off S-IVB 9566.2 T7 +15.4 CCSCommand

a n d C ryo O n

Water Coolant Valve Closed IU 15,213.7 T7 +5662.9 LVDCFunction

LHp Tank Continuous Vent Valve S-IVB 16,953.4 T8 +17.5 CCSCommand-Close On

LH? Tank Continuous Vent Valve S-IVB 16,956.2 T8 +20.3 CCS Command-Close Off

LH9 Tank Repress Control Valve S-IVB 17,096.6 T8 +160.8 CCS Command-Open Off

LH Tank Repress Control Valve S-IVB 17,098.5 T8 +162.7 CCS Command

20 pen O ff

Ambient Repress Mode Selector On S-IVB 17,100.0 T8 +164.1 CCSCommand

an d C ryo O ff

S-IVB Ullage Engine No. I On S-IVB 17,136.5 T8 +200.6 CCSCommand

S-IVB Ullage Engine No. 2 On S-IVB 17,137,8 TB +202.0 CCSCommand

LOXTank Repress Control Valve S-IVB 17,15_.I T8 +217.3 CCSCommandO pen O n

Auxiliary Hydraulic PumpFlight S-IVB 17,186.4 T8 +250.6 CCSCommandM ode O n

Passivation Enable S-IVB 17,299.8 T8 +364.0 CCSCommand

Engine Helium Control Valve Open On S-IVB 17,300.6 T8 +364.8 CCS Command

Engine Mainstage Control Valve Open S-IVB 17,301.3 T8 +365.5 CCSCommandO n

Engine Mainstage Control Valve Open S-IVB 17,310.3 T8 +374.5 CCSCommandO ff

Prevalves Close On S-IVB 17,311.4 T8 +375,5 CCSCommand

Engine Helium Control Valve Open S-IVB 17,312.2 T8 +376.3 CCSCommandO ff

Engine Helium Control Valve Open S-IVB 17,324.7 T8 +352.9 CCSCommandO n

Engine Helium Control Valve Open S-IVB 17,336.3 T8 +400.4 CCSCommandO ff

Passivation Disable S-IVB 17,337.0 T8 +401.2 CCSCommand

LOXTank Repress Control Valve Open S-IVB 17,355.5 T8 +419.6 CCSCommandO ff

LH2 Tank Repress Control Valve Open S-IV8 17,356.9 T8 +421.0 CCSCommandO n

Prevalves Close Off S-IVB 17,358.3 T8 +422.5 CCSCommand

Ambient Repress Mode Selector Off S-IV8 17,386.3 T8 +450.4 CCSCommand

an d C ryo O n

2-I0



T able 2-3. Variable T im e and Com manded S witch S elector E vents (Continued)

FUNC'flON STAGE RANGETIME TIMEFROMBASE(SEC) (SEC) REMARKS

Passivation Enable S-IV8 17,408.5. T8 +472.6 CCS Command

Engine IgnitionPhase Control Valve S-IVB 17,409.9 T8 +474.1 CCS CommandOpen On

EngineHelium Control Valve Open On S-IVB 17,411.3 T8 +475.5 CCS Command

S-IVB UllageEngineNo. I Off S-IVB 17,415.6 T8 +479.7 CCS Command

S-IVB UllageEngineNo. 2 Off S-IVB ]7,417.0 T8 +481.I CCS Command

Engine IgnitionPhase Control Valve S-IVB 17,458.9 T8 +523.0 CCS CommandOpen Off

Engine Helium Control Valve Open Off S-IVB 17,460.3 T8 +524.4 CES Command

LH2 Tank Continuous Vent Orifice S-IV8 17,496.1 T8 +560.2 CCS CommandShutoff Valve Open On

LHp Tank ContinuousVent Relief S-IVB 17,497.5 T8 +561.6 CCS Command-Override Shutoff Valve Open On

LH2 Tank ContinuousVent Orifice S-IVB 17,500.3 T8 +564.5 CCS CommandShutoff Valve Open On

LH? Tank ContinuousVent Relief S-IVB 17,501.8 T8 +565.9 CCS Command-Override Shutoff Valve Open Off

WaterCoolantValveOpen IU 17,919.4 T8 +983.5 LVDCFunction

Water CoolantValveClosed IU 18,219.5 T8 +1283.6 LVDCFunction

S-IVB UllageEngine No. I Off S-[V8 19,743.3 T8 +2807.4 CCS Command

S-IVB UllageEngine No. 2 Off S-IV8 19,744.9 T8 +2809.0 CCS Command

AuxiliaryHydraulicPump Flight S-IV8 ]9,792.3 T8 +2856.4 CCS CommandMode On

AuxiliaryHydraulicPump Flight S-IV8 19,928.1 T8 +2992.3 CCS CommandMode Off

WaterCoolantValveClosed IU 23,329.3 T8 +6393.4 LVDCFunction

PCM Coax Switch High Gain Antenna IU 24,053.8 T8 +7118.0 CCS Command

CCS Coax Switch Fail Safe and High IU 24,053.9 T8 +7118.1 CCS CommandG ai n A nt en na

PCM Coax Switch High Gain Antenna IU 24,054.0 T8 +7118.2 CCS Command

CCS Coax Switch Fail Safe and High IU 24,054.1 T8 +7118.2 CCS CommandG ai n A nt en na

WaterCoolantValveOpen IU 25,133.1 T8 +8197.2 LVDCFunction

Water CoolantValveClosed IU 25,433.8 T8 +8497.9 LVDCFunction

2 - 1 1



T able 2-3 a lso con ta ins the tim es of in itia tion of the specia l sequence

of switch selecto r events wh ich were program ed to be in itia ted by tele-

metry station acquisition and included the following calibration sequence:

FUNCTION STAGE TIME(SEC)

In-Flight Calibration ModeOn S-IVB Acquisition +60.0

Telemetry Calibrator In-Flight IU Acquisition +60.2Calibrator O n

TMCalibrate On S-IVB Acquisition +60.4

Telemetry Calibrator In-Flight IU Acquisition +65.2Calibrate O n

TMCalibrate Off S-IVB Acquisition +65.4

In-Flight Calibration Mode Off S-IVB Acquisition +66.0

In addition , known ground com m ands sent to the L VDC are included inthis table.

2-12



S E CT IO N 3

L A UN C H O P E R AT IO N S

3.1 S UM M A R Y

The ground systems supporting the AS-505/Apollo I0 countdown and launchperformed exceptionally well. There were no significant fa ilures oranomalies. S everal systems experienced component failures andmalfunctions, but these problem s did not cause any holds or significantdelays in the scheduled sequences of launch operations. L aunch occurred

at 1 2:49 :0 0 E astern Daylight T im e (E DT ), M ay 1 8, 1 9 6 9 .

T he A po llo I0 vehicle was the first to be launched from P ad 39 B of the

S aturn com plex. Dam age to the pad, L aunch Um bilical T ower (L UT ) andsupport equipm ent from the blast and flam e im pingem ent was minor. Ahydraulic o il fire occurred in S ervice A rm (S A ) N o. 1 contro l console.

3.2 P RE LA UN CH M ILE STO N ES

A chronological summary of events and preparations leading to the launchof the A S -5D5 is conta ined in T able 3-I.

3.3 CO UN TDO W NE VE NT S

The AS-505/Apollo I0 terminal countdown was picked up at 21:00:00 EDT,M ay 1 6 , 1 96 9 . T he countdown proceeded as planned with no unscheduled

holds. T he schedu led l-hour hold at -3 hours 30 m inutes in the count wasutilized to overcome a slight delay in propellant loading. L aunch

occurred at 1 2:49 :0 0 E DT , M ay 1 8, 1 969 .

3.4 P RO PE LL A NT L O ADIN G

3.4.1 R P -I L oading

T he R P -I system successfully supported the launch countdown. During theautomatic replenish operations at approximately -12 hours, the fast fillvalve "open" indication dropped out causing system shutdown. R eplenishoperations were rein itia ted in the manual_ mode and were completed

satisfactorily. The problem was subsequently traced to improperlyadjusted fast fill valve limit switches. A lthough attempts at readjust-ment were unsuccessful, there was no significant impact on remaining R P -I

operations. T he fast fill valve is not used during the countdown after

3 - I



T able 3-I. A S -50 5 P relaunch M ilestones

DATE ACTIVITYOREVENTi , I

October 16, 1968 LM-4 Arrival

November 23, 1968 CSM-I06 Arrival

November 27, 1968 S-IC Stage Arrival

December 3, 1968 S-II Stage Arrival

December I0, 1968 S-IVB Stage Arrival

December15, 1968 IU Arrival

December 30, 1968 Launch Vehicle (LV) Erection Completed

January 17, 1969 Final MannedAltitude Run

February 3, 1969 LV Propellant Dispersion/MalfunctionO vera ll T est (O A T )

February 6, 1969 Spacecraft (SC) Erection

February 27, 1969 Space Vehicle (SV) Electrical Mate

March5, 1969 SVOAT(plugs in)

March II, 1969 SVTransfer to Pad 39B

March 28, 1969 Emergency Egress Test Completed

April 19, 1969 SV Flight Readiness Test (FRT) Completed

April 25, 1969 SV Hypergol Load Completed

May2, 1969 S-IC RP-I Loaded

May5, 1969 CDDTWet) Completed

May6, 1969 CDDT(Dry) Completed

May 18, 1969 SV Launched on Schedule

rep len ish is com ple ted. If unscheduled rep len ish had been requ ired, it

cou ld have been acco m plished, as before, in the m anua! m ode.

During the autom atic R P -I level adjust a t abou t -50 m inu tes, an anom aly

occurred which cau sed the level a djust va lve to close sligh tly la te.A s a resu lt the R P -I fligh t m ass percen tag e (w h ich can no rm a lly be

adjusted to I0 0 ±0 .0 2 percent) was adjusted to 9 9 .81 -percen t, bu t was

still w ith in L aunch M ission R ules requ irem ents o f I0 0 ± 0 .2 percen t. T he

sam e problem occurred durin g C ountdown D em onstra tion T est (C DD T ), how-

ever post CDDT troubleshooting revealed no problem s with the P ropellan t

T anking C ontro l S ystem (P T C S ). F urther postlaunch investiga tion has

iso la ted the problem to a defective printed circu it card in the P T C S .

3-2





3.5.2 Forward B ulkhead Uninsulated A rea

The circuit in let pressure fR r the forward bulkhead uninsulated arearemained steady at 1.14 N/cm_ (1.65 psig) through propellant loading

and until the tim e of launch. T he outlet p ressure was at or near zerothroughout the countdown. N itrogen contamination exceeded one percentat the beginning of L O X load and diminished to an insignificant level

after L H2 loading.

3.5.3 L H2 T ank S idewall

The sidewall inlet pressure remained steady between 1.04 and 1.20 N/cm2(I.51 and 1.74 psig). The outlet pressure decayed from 1.09 to 0.72 N/c_2(I 58 to 1.05 psig) during LOX tanking, and further decayed to 0.41N/cmZ(0 .6 psig) during L H2 loadin_. A t the tim e of launch the outlet pressurehad recovered to 0.49 N/cm_ (0.71 psig). Hydrogen and nitrogen contam-ination exceeded one percent about 2 hours before launch, but as the

time of launch approached, all sidewall contamination readings becameinsignificant.

Insulation outer surface temperatures on the sidewall were warmer than onany preceeding stage. M inimum temperature recorded during the countdownwas 26 9 °K (25°F ). O n S -II-50 4, the lowest tem perature experienced was

21 9 °K (-6 5°F ). T he absence of frost was noteworthy since all targetsand other markings were clearly visible.

3.5.4 Bolting/J-Ring

The bolting ring inlet pressure diminished steadily from 1.19 to 0.98 N/cm2

(1.73 to 1.42 psig) at launch. The outlet pressure decayed from 0.9 N/cm_(I.3 psig) to sligh tly below zero at the tim e of L H2 loading , thenrem ained at or below zero until launch. HazardO us gas concentration

readings were questionable due to problems experienced with the groundanalyzer system.

3.5.5 F eedline E lbow

T he feedline elbow circuit in let pressure remained S teady between 2.28and 2.48 N/cm2 (3.3 and 3.6 psig). Outlet pressure varied from 0.71 to1.2 N/cm2 (1.04 to 1.74 psig) during tanking but otherwise remainedsteady until launch. There were no significant contamiDation readings.

3.5.6 Common B ulkhead

P urge pressures remained approximately steady during the period thebulkhead was purged. E vacuation beqan at -3 hours 24 minutes and

pressures decreased below 2.07 N/cm_ (3 psia) within 50 minutes. Thisis well within acceptable limits. T here were no significant hazardousgas readings in this purge circuit.

3 - 4





3.6 .2.2 S -II S tage GS E . T he S -ll stage GS E perform ed satisfactorilyduring countdown and launch. B last damage to the mechanical and electri-cal equipm ent was very m inor. During the postlaunch inspection a L O Xumbilical debris va lve was found to be open, although the proper closed

signal was received at launch. T here were no significant system failures

or anomalies, and all minor discrepancieswere promptly corrected duringthe countdown. O ne of the m inor discrepancies was leakage of a pressure

regulator in the pneum atic supply console at -22 hours 30 m inutes. T h isregulator was replaced. A nother regulator in th is console requiredadjustment at -8 hours 25 minutes.

3.6 .2.3 S -IVB S tage GS E . A ll S -IVB GS E systems operated satisfactorilyduring the countdown. T he only problem reported was a faulty connectionin the pneum atic console distributor cable. N o m ajor damage was found

during the postlaunch inspection.

3.6 .2.4 IU S tage G S E . T he IU G S E perform ance during countdown was

satisfactory. N o anomalies were encountered with the mechanical equip-ment. S everal malfunctions and anomalies occurred in the electrical GS E

during launch preparations and countdown which were promptly correctedto maintain launch readiness. T he systems involved were the Ground

Computer, Integration N etworks, Digital Data A cquisition S ystem,S tabilizer, Count Clock and DE E -6 .

S upport of the D E E -6 was lost after lifto ff. Due to h igh discreteactivity encountered at lifto ff, it is norm al for the DE E -6 to runbehind in the processing of backlogged discrete events. A s a result,the last indication of data output occurred at -1 .6 3 seconds. P ost-

launch inspection indicated a possibility of a memory parity halt as thereason for DE E -6 fa ilure. Through software manipulation, the liftoff

data were obtained from the backlogged information and transmitted tothe master printer.

3.6 .3 GS E Camera Coverage

O n review of the film coverage of the GS E at launch the fo llowinganomalies were noted:

a. S -II S tage F orward S A CS A N o. 5) electrica l um bilica l access panelwas observed not secured and flapping after liftoff.

b. T he retracting cable on the S erv ice M odu le S A (S A N o. 8) did no tretract the umbilical carrier to the boom after disconnection and

was observed hanging 1 .2 to 1 .5 m eters (4 to 5 ft) after fu llretraction of the arm.

c. A G S E cabine t on the east s ide 1 8 .3 m eter (6 0 ft) leve l o f the L U Twas observed on fire,and the cabinet doors open after S -IC flam e

impingement. F lame im pingement obscured time of door opening. L astappearance of cabinet in tact was at 8 seconds and the doors wereobserved open at 1 7 seconds.

3 - 6



S ECT IO N 4

TRAJECTORY

4.1 S UM M A R Y

T he trajecto_ parameters from launch to translunar injection were closeto nom inal. T he vehicle was launched on an azim uth 9 0 degrees east ofnorth . A ro ll m aneuver was in itia ted at 1 3.0 5 seconds range tim e that

placed the w_hicle on a flight azimuth of 7 2.0 28 degrees east of north .

T he space-fixed velocity at S -IC O utboard E ngine Cutoff (O E CO ) was10.81 m/s (35.47 ft/s) greater than nominal. The space-fixed velocityat S-II OECOwas 13.22 m/s (43.37 ft/s) less than nominal, The space-

fixed velocity at S-IVB first guidance cutoff was 0.07 m/s (0.23 ft/s)greater than nominal. The altitude at S-IVB first guidance cutoff was0 .0 3 k ilom eter (0 .0 1 n m i) lower than nom ina l and the surface range was

0 .9 2 kilom eter (0 .50 n m i) greater than nom inal.

The space-fixedvelocityat parkingorbit insertionwas 0.07 m/s

(0.23ft/s) "lessthan nominaland the flightpath angle was 0.0059degreeless than nominal. T he eccentricitywas 0.000037 greater than nominal.

The apogee and perigee were 0.13 kilometer (0.07 n mi) and 0.62 kilometer

(0.33 n mi) lower than nominal, respectively.

The parameters at translunar injection were also very close to nominal.

The eccentricitywas 0.00002 less than nominal, the inclination was 0.007

degree greater than nominal, the node was 0.022 degree lower than nominaland C3 was 868 m2/s2 (9345ft2/s2) less thannominal. The total space-

fixed velocitywas 2.39 m/s (7.84ft/s) less thannominaland the altitudewas 2.87 kilometers (1.55 n mi) higher than nominal.

The actual impact locations for the spent S-IC and S-II stages were deter-

mined by a theoretical free-flight simulation. The surface range for the

S-IC impact point was 4.19 kilometers (2.26 n mi) less than nominal. Thesurface range for the S-II impact point was 34.57 kilometers (18.67 n mi)

less than nominal.

4.2 TR ACKIN G DA TA UT IL IZA TIO N

4,2.1 T racking During the A scent P hase of F light

T racking data were obtained during the period from the time of first m otionthrough parking orbit insertion.

4 -I





30 0 0 - 1 40 -,

- ACTUAL'. ....N(,MINAL /

_7S-IC OECO /

2500. !00-1 %_7S-ITOECO

_S-IVBFIRSTECO I''"

2000- i60-1 j_

/! ALTITUDE- ---_

1500. A / /

/ //'g. SURF CE RANGE ,/

/ / /ooo- 8o...I /,

/ // / / ---- CROSSANGE

- j5oo- 4o-I / y

O- O" _ _ _7

0 100. 200 300 400 500 600 700 800

R AN GE T IME , S EC ON DS

Figure 4-I. Ascent Trajectory Position Comparison

Mach number and dynamic pressure are shown in Figure 4-4. These parameters

were calculatedusingmeteorologicaldata measuredto an altitudeof 89.75

kilometers (46.46 n mi). Above this altitude the measured data were

merged into the U. S. Standard Reference Atmosphere.

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, re spe cti vel y.

The free-flight trajectories of the spent S-IC and S-IT stages were simu-

lated 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 presentsa comparisonof nominaland free-flight

parameters 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 andS-II stages.

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

4 - 3



3 2

I iI" ACTU_

--- N_INAL

2e / _ s-lcOECOs-IfEco /_ s-,v,,RSTCO/

24-

\ /"

20"

/

16-g

_ //--_ SPACE-FIXED_ _ VELOCITY

L8-

/ _IGHT PATHNGLE

,- _i \ ,/

°" /

0 lO0 200 300 400 500 600 7_ _0

_GE TIME, SECONDS

F igure 4-2. A scent T rajecto_ S pace-F ixed Velocityand F light P ath A ngle Comparisons

±250 meters (±820 ft) in position components and ±0.7 m/s (±2.3 ft/s) invelocity components. The actual and nominal parking orbit insertion

parameters are presented in T able 4-5. T he ground track from insertionto S-IVB/CSM separation is given in Figure 4-5.

4 - 4



4 .

o_

_ 2-_-_J

0 IO0 200 300 400 500 600 700 800

RANGETIME,SECONDS

Figure 4-3. Ascent Trajectory Acceleration Comparison

4.3.3 Injection T ra jectory

Comparisons between the actual and nominal total space-fixed velocity andflight path angle are shown in Figure 4-6. The actual and nominal total

inertial acceleration comparisons are presented in Figure 4-7. Through-out the S-IVB second burn phase of flight, the space-fixed velocity andthe flight path angle were very close to nominal with deviations morenoticeable towards the end of the time period.

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

4-5



4.0

3.5

/_ _ ACTUAL--DYNAMICPRESSURE- , k --- NOMINAL

l . O

_., -- MACHNUMBER _,

0 ,_ _....__ ..._,._. ] x_.0 20 40 60 80 lO0 120 140 ]60

RANGE TIME, SECONDS

Figure 4-4. Dynamic Pressure and Mach Number Comparisons

4.3.4 Post TLI Trajectory

T he post T ranslunar Injection (T LI) trajectory spans the time interval from

TLI to S-IVB/CSMseparation. The postTLI trajectorywas obtainedby inte-

grating the translunar injection conditions, derived from the injection

trajectorysolution,to S-IVB/CSMseparation. A comparisonof S-IVB/CSM

separation conditions is presented in Table 4-3. The post TLI trackingdata were received and were used to verify the post TLI trajectory.

4 - 6



T able 4-I. Comparison of S ignificant T rajectory E vents

EVENT PARAMETER ACTUAL NOMINAL ACT-NOM

First Motion Range Time, sec 0.25 0.25 D.O

Total Inertial Acceleration, mZs2 10.40 10.92 -0.52(ft/s _} (34.12) (35.83) (-1.71)

Mach 1 Range Time, sec 66.8 65.9 0.9

Altitude, km 7.86 7.74 0.12

(nmi) (4.24) (4.18) (0.06)

Maximum Dynamic Range Time, sec 82.6 8l.l 1.5Pressure

Dynamic Pressure, N/c_2 3.324 3.384 -0.060(Ibf/ftz) (694.2) (706.8) (-12.6)

Altitude, km 13.22 12.91 0.31

(n mi) (7.14) (6.97) (0.17)

Maximum TotalInertlal

Acceleration: S-IC Range Time, sec 161.71 160.16 1.55

Acceleration, m/s 2 38.47 38.01 0.46(ft/s2) {126.21) (124.70) (1.51)

S-IT Range Time, sec 460.69 459.28 1.41

Acceleratlon, m/s 2 17.82 17.75 0.07(ft/s 2) (58.46) (58.23) (0.23)

S-IVB Ist Burn Range Time, sec 703.84 703.56 0.28

Acceleration, m/s 2 6.89 6.85 0.04

(ft/s 2) (22.60) (22.47) (0.13)

S-IVB2nd Burn RangeTime, sec 9,550.66 9,548.67 1.99

Acceleration, m/s 2 14.60 14.63 -0.03

(ft/s2) (47.90) (48.00) (-0.I0)

Ma xi mu m E ar th. -F ix ed

Veloclty: S-IC RangeTime,sec 161.96 160.91 1.05

Velocity, m/s 2,388.34 2,380.96 7.38(ft/s) (7,835.76) (7,811.55) (24.21)

S-II RangeTime,sec 553.50 555.04 -I.54

Velocity, m/s 6,497.67 6,511.84 -14.17(ft/s) (21,317.81) (21,364.30) i(-46.49)

S-IVB IstBurn RangeTime,sec 713.76 713.48 0.28

Velocity, m/s 7,389.65 7,389.70 -0.05

(ft/s) (24,244.26) (24,244.42) (-0.16)

S-IVB 2nd Burn Range Time, sec 9,551.30 9,549.25 2.05Velocity, m/s 10,439.91 10,442.07 -2.16

(ft/s) (34,251.67) (34,258.76) (-7.09)

Apex: S-ICStage RangeTime,sec 266.87 267.88 -1.01

Altitude, km I12.25 I15.24 -2.99(n mi) (60.61) (62.22) (-1.61)

Surface Range, km 320.21 321.45 -1.24(n mi) (172.90) (173.57) (-0.67)

S-ll Stage Range Time, sec 597.21 596.33 0.88

Altitude, km 189.48 190.25 -O.ll

(n mi) (102.31) (I02.73) (-0.42)

Surface Range, km 1,916.93 1,912.52 4.41(n mi) (I,035.06) (1,032.68) (2.38)

4 - 7



T able 4-2. Comparison of Cutoff E vents

AC3UAL NOMINAL ACT-NON

ACTUAL NOMINAL ACT-NONARAME,ER 1 1.... 1.....S-IC CECO (ENGINE SOLENOID) S-IC OECO (ENGINE SOLENOID)

--Range Time, sec 135,161 ]35.26 -O.IO 161.63 160,20 1.43

Altitude, km 43.39 _ 44.56 -1.17, 85.28 65.79 -0,5l(n mi) (23.43) (24.06) (-0.63)I (35.25) (35.52) (-0,27)

Surface Range, km 46.32 46.88 -0.56 93.38 91.33 2.05(n mi) (25.01} (25.3]) (-0.30) (50.42) (49.31) (1.l])

Space Fixed Velocity. m/s 1,973.03 2,001.52 -28.49 2,751.91 2,741,10 lO,BI(ft/s) (6,473.20) (6,566.67) (-93.47) (9.028.58) (8,993.I1) (35.47)

Hight Path Angle, dog 22.807 23.1531 -0.346 ]8.946 19.545 I -0.599

Heading Angle, dog 76.48] 76.217 0.244 75.538 75.360 0.178

Cross Range, km 0.23 0.171 8.06 0.60 i 0.33 0.27(nmi) (0.]2) (0.09) (0.03) (0.32) ' (0.18) (O.14)

Cross Range Velocity, m/s 10.49 4.21 6.28 17.09 9.08 8.8I(ft/s) (34.42) (13.81) (20,61) (58.69) (29.79) (28.90)

S-I] CECO (ENGINE SOLENO_]D) __..S-IT OECO (ENGINE SOLENOID)

Range Time, sec 460,61 459.21 1.40 552.64 554.13 -I.49

179.00 180.43 -1.43 187.43 108.32 -0.89

AItitude'(_mmi ) (96.85)i (97.42) (-0.77) (]01.20) (10].68) (-0.48)

SurfaceRange(nkm 1,109.50 I,I02.86 6.64 ],836.56 ],646.50 -9.94mi) (599.08) (595.50) (3.58) (883.67) (889.04) (-5.37)

Space-Fixed Velocity, m/s 5,678.47, 5,667.07 - ]I.40 6,898.24 6,911.46 -13.22(ft/s) (18,630.15)!(18,592.75) (37.40) (22,632.02) (22.615.39) (-43.37)

FlightPathAngle,dog 1.029 0.864 0.I65 0.741 0.735 0.006

Heading Angle, dog i 79.585 79.612 -0.027 82.458 82.844 -0.006

]5.89 15.02 O.B7 28.68 28.76 -0.08

Cross Range,(_mmi ) (8.58) (8.11) (0.47) (]5.49) (15,53) (-0.04)

Cross Range Velocity, m/s ! I09.59 ]14.99 -5.40 172.16 I76.64 -4.48(ft/s) (359.55} (377,26) (-]7.71) (564.83) (579.53) (-]4,70)

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

703.76 703.48 0.28 9,550.58 9,848.64 ].94Range Time, sec

19].47 191.50 -0.03 319,81 317,02 2.79Al'titude'(_mmi) (103.39) (I03.40) (-0.01) (172.68) (171,18) (1.50)

Surface Range, km 2,650.2] 2,649.29 0.92(n mi) '1,431.00) (1,430.50) (0.50)

Space-Fixed Velocity, m/s 7,79].42 7 791.35 0.07 ]0,846.56 I0,849,I2 -2.56

(ft/s) (25,582.40 (25,582.]7) (0.23 (35,585.83) (35,594.23) (-8.40)

Flight Path Angle. deg -0.0064 -0.0002 -0.0062 6.927 6.867 0.060

Heading Angle, dog 88.497 88.483 0,014 61.258 81.30] -0.043

Cross Range, km 62.10 62.12 -0.02

(n mi) (33.53) (33.54) (-O,Ol)

Cross Range Velocity, m/s 275.31 274.12 1,19(ft/s) (903.25) (899.34) (3,9])

Eccentricity 0.97688 0.97698 -0.0001(

-1,396,438 -1,390,603 -5833

C3'' m2/s2 _ (-15.031,112)(-14,968,326) (-62,786)(ft2/s _)

Inclination, dog 31.701 31.693 0.008

Descendingode,deg I23,51] 123.536 -0.025

* C3 is twice the specific energy of orbit

C3 V2 2p

where V = Inertial Velocity= Gravitational Constant

Radius vector from center of earth

4 - 8



T able 4-3. C om parison of S epara tion E vents

ACTUAL NOMINAL ACT-NOMPARAMETER

S-IO'S-II SEPARATION

RangeTime,sec 162.31 160.91 1.40

Altitude,km 65.89 66.43 -0.54

(n mi) (35.58) (35.87) (-0.29)

SurfaceRange,km 94.88 92.85 2.03

(n mi) (51.231 (50.13) (I.101

Space-Fixed Velocity, m/s 2,759.29 2,750.70 8.59

(ft/s) (9,052.79) (9,024.61) (28.181

Flight Path Angle, deg 18.848 19.444 -0.596

HeadingAngle,deg 75.538 15.355 0.183

CrossRange,km 0.61 0.33 0.28

(n mi) (0.33) (0.181 (0.151

Cross Range Velocity, m/s 18.05 9.20 8.B5

(ft/s) (59.22) (30.181 (29.04)

Geodetic Latitude, deg N 28.883 28.879 0.004

Longitude, deg E -79.694 -79.714 D.020

S -IIIS -[VB S EPA RA TIO N

Range Time, sec 553.50 555.04 -I.54

Altitude,km 187.51 188.40 -0.89

(n mi) (I01.251 (I01.73) (-0.48)

Surface Range, km 1,642.05 1,652.19 -10.14

(n mi) (886.64) (892.11) (-5.47)

Space-Fixed Velocity m/s 6,900.65 6,914.90 -14.25

(ft/s) (22,639.93) (22,686.68) (-46.75)

Flight Path Angle, deg 0.730 0.725 0.005

HeadingAngle,deg 82.490 82.577 -0.087

Cross Range, km 28.83 28.92 -0.09

(nmi) (15.571 (15.62) (-0.05)

Cross Range Velocity, m/s 172.65 177.13 -4.48

(ft/s) (566.44) (581.14) (-14.701

Geodetic Latitude, deg N 31.925 31.939 -0.014

Longitude, deg E -63.965 -63.858 -0.107

S-[VB/CSM SEPARATION

Range Time, sec 10,962.4 II,004.9 -42.5

Altitude, km 6,486.86 6,722.07 -235.21

(n mi) (3,502.63) (3,629.63) (-127.001

Space-Fixed Velocity, m/s 7,787.25 7,7]5.38 71.87

(ft/s) (25,548.72) (25,312.93) (235.79)

F]ightPath Angle,deg 43.93 44.45 -0.52

Heading Angle, deg 67.47 67.88 -0.41


Longitude,deg E -139.826 -138.933 -0.893

4-9



T able 4-4. S tage Im pact L oca tion

PARAMETER ACTUAL NOMINAL ACT-NOM

S -IC S T A G E IM P A C T

RangeTime, sec 539.12 542.07 -2.95

Surface Range, km 645.98 650.17 -4.19

(n mi) (348.80) (351.06) (-2.26)

Cross Range, km 9.96 7.69 2.27

(nmi) (5.38) (4.15) (1.23)


Longitude, deg E -74.207 -74.172 -0.035

S -II S T A G E IM P A C T

Range Time, sec 1,217.89 1,222.49 -4.60

Surface Range, km 4,424.97 4,459.54 -34.57

(n mi) (2,389.29) 2,407.96) (-18.67)

Cross Range, km 144.35 147.44 -3.09(n mi) (77.94) (79.61) (-1.67)

Geodetic Latitude, deg N 31.522 31.457 0.065

Longitude, deg E -34.512 -34.158 -0.354

4.3.5 S-IVB/IU Post Separation Trajectory

The S-IVB/IU was placed on a lunar slingshot trajectory close to nominal.T his was accom plished by orien ting the stage in a retrograde altitude

(p itch = 1 9 4 degrees w ith respect to loca l h orizon ta l, ya w = 0 degree ,

ro ll : 1 89 degrees) and app lying a velocity increase a long the positive

X body axis. T he velocity increase w as derived from a com bina tion o f L O Xdump, LH2 vent, Auxiliary Propulsion System (APS) burn and JL2 enginepropellant lead experim ent. T he engine propellant lead experimentconsisted o f a 27 3-second A P S burn , a 9 -second L O X lead and a 53-second

L H2 lea d. T he fina l A P S burn w as sho rten ed in rea l tim e from 1 55 seconds

to approxim ate ly 8 seconds based on updated L O X residuals wh ich were no t

considered at the tim e prefligh t slingshot targeting was perform ed. A

tim e h istory of the long itudina l velocity increase subsequentto T im e

4-10



Table 4-5. ParkingOrbit InsertionConditions

PARAMETER ACTUAL NOMINAL ACT-NOM

RangeTime,sec 713.76 713.48 0.28

Altitude,m 191.37 191.51 -0o14

(nmi) (I03.33) (I03.41) (-0.08)

Space-FixedVelocity,m/s 7,793.09 7,793.16 -0.07

(ft/s) (25,567.88) (25,568.11) (-0.23)

FlightPathAngle,deg -0.0049 O.OOlO -0.0059

Headingngle,deg 88.933 88.918 0.015

Inclination,eg 32.546 32.545 O.OOl

Descendingode,deg 123.132 123.148 -0.016

Eccentricity 0.000086 0.000049 0.000037

Apogee*,m 185.79 185.92 -0.13

(nmi) (I00.32) (I00.39) (-0.07)

Perigee*,m 184.66 185.28 -0.62

(nmi) (99.71) (100.04)(-0.33)

Period,in 88.20 88.20 0.00

GeodeticLatitude,degN 32.700 32.699 O.OOl

Longitude,egE -52.526 -52.537 O.Ol]

*Based on a spherical earth of radius 6378.165 km (3443.934 n mi).

Base 8 (T8) is presented in Figure 4-8 and Table 4-7 lists the velocitygained during the various portions of slingshot maneuver.

The S-IVB/IUclosestapproachof 3112kilometers_(1680n mi) above thelunar surface occurred at 78.851 hours into the mission. The actual and

nominal conditions at closest approach are presented in Table 4-8. Thevelocityof the S-IVB/IUrelativeto earth is presentedin Figure4-9.T his illustrates how the influence of the moon's gravity imparted energy

to the S-IVB/ILI.

Some of the heliocentricorbit parametersof the S-IVB/IUare presented

in Table 4-9. The same parameters for the earth's orbit are also listed

for comparison.

4 - I I



0 F IR S T R E V O L U T IO N

Q S ECO N DR EV O LU T IO N

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R ES TA R T SEP A R A T:dNINSERTION CO_A,=D ,

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L ON GIT UDE , deg

F igure 4-5 . Ground Track



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I-- ACTUAL LI I. .. . N O MI NA L

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-2 O 00 I00 200 300 400 500 600 700 800 900 I000 0 1O0 200 300 400 500 600 700 800 900 1000

TIMEFROM6, SECONDS TIMEFROM6, SECONDS

02:24:00 02:27:00 02:30:00 02:33:00 02:36:00 02:39:000-2:24:0002:27:00 02:30:00 02:33:00 02:36:00 02:39:00 RANGETIME, HOURS:MINUTES:SECONDS

RANGE TIME, HOURS:MINUTES:SECONDS

Figure 4-6. Injection Phase Space-Fixed Velocity Figure 4-7. Injection Phase Acceleration Comparison

and Flight Path Angle Comparisons





42 I"" CVS* "

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12 / *ContinuousVent System (CVS) _

/continues to 2700 seconds.

At 2800 seconds an 8 second

10 APSburnoccurreditha total

accumulated velocity change at

8 2810 seconds of 44.2 m/s (145.0 ft/s).

4

/00 200 400 600 800 1000 1200 1400

TIME FROM T 8 , SECONDS

I I I I I I I I

4:42:1 6 4::45 :36 4:48:56 4:52:1 6 4:55 :36 4:58 :56 5 :0 2:1 6 5:0 5 :36

RANGETIME, HOURS:MINUTES:SECONDS

Figure 4-8. Velocity Increments Due to Slingshot Activity

4 - 1 5



Table 4-8. L unar Close A pproach P arameters

PARAMETER UNITS ACTUAL NOMINAL

LunarRadiusof km 4850 4748

ClosestApproach (n mi) (2619) (2564)

Altitude Above km 3112 3010

LunarSurface (nmi) (1680) (1625)

TimefromLaunch hr 78.9 78.5

Velocity Increase km/s 0.850 0.861Relative to Earth, (n mi/s) (0.459) (0.465)Due to Lunar Influence

Tabl_ 4-9 . Heliocentric O rbit P arameters

PARAMETER UNITS S-IVB/IU EARTH

Semi-majorAxis km 1.4398 x 108 1.4900 x I0'_(n mi) (0.7774 x 108) (0.8045 x 1081

Aphelion km 1.5216x 108 1.5115x 108(n mi) (0.8216 x 108) (0.8161 x 108)

Perihelion km 1.3581 x 108 1.4684x 108

(n mi) (0.7333 x 108) (0.7929 x 108)

Inclination deg 23.46 23.44

Period days 344.88 365.25

4-16



2.8 q

4TH DAY

HR I 5TH DAY2.4

6TH DAY IOTH DAY 21ST DAYIII

2.0 IIII

l

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0. 8

0.4"

O '

0 0,2 0,4 0,6 0.8 1.0 1,2 1.4 1.6 1.8

DISTANCE(EARTH-VEHICLE), 106 KM

Figure 4-9, S-IVB/IU Velocity Relative to Earth Distance

4-17/4-18



S E CT IO N 5

S -IC P RO PUL S IO N

5.1 S U M M A R Y

Unless otherwise stated, all predicted propulsion performance parametersused in th is section are based on the la test p relaunch S -IC propu ls ion

perform ance prediction , which was not incorpora ted in the L aunch V ehic le

O perationa l F ligh t T ra jectory, dated A pril 1 7 , 1 9 6 9 . T he principa l change

in the latest S-IC propulsion prediction was in the predicted thrust levelsof the five F-I engines. This amounted to decreasing the thrusts used inthe earlier predictions by approximately 40,000 Newtons (9000 Ibf) perengine.

A ll S -IC propu ls ion system s perform ed satisfacto rily. A t the 35 to 38-

second tim e slice, average eng ine thrust reduced to sta ndard conditions

was 0.20 percent lower than predicted. Average reduced specific impulsewas 0 .0 3 percent lower than predicted, and reduced p ro pellan t consum ption

ra te was 0 .1 58 percent lower than predicted. A lthough the average th rust

dev ia tion fro m p red icted w as sm a ll, en g ine N o . 1 d id run a t a leve l o f6 ,6 1 1 ,0 0 0 N ew to ns (1 ,486 ,0 0 0 Ibf) a t th e 35 to 38-seco nd tim e slice , w h ich

was sign ificantly lower than th e predic ted level o f 6 ,7 0 8,0 0 0 N ewtons

(I ,50 8,0 0 0 Ibf).

C ente r E ng ine C u to ff (C E C O ) w a s in itia ted by the Instrum en t U n it (IU ) a t

1 35.1 6 seconds as p lanned. O utboard E ng ine C uto ff (O E C O ), in itia ted byLOX low level sensors, occurred at 161.63 seconds which was 1.43 seconds

la ter than predicted in the F ligh t T ra jecto ry. H owever, based on thelatest S-IC propulsion prediction, OECO occurred only 0.63 second laterthan predicted. T h is is a sm all d ifference com pared to the predicted 3

sigm a lim its o f ± 7 .0 5 seco nds. T he L O X res idua l a t O E C O w a s 1 8,41 2 kilo -

g ram s (40 ,59 2 Ibm ) com p ared to th e predic ted 1 7 ,57 9 k ilog ra m s (38 ,7 56 Ibm ).T h e fu e l res idua l a t O E C O w a s 1 2,9 44 k ilog ra m s (28,537 Ibm ) co m pared to

the predicted 1 6 ,0 29 k ilogram s (35,338 Ibm ).

5.2 S -IC IG N IT IO N T R A N S IE N T P E R F O R M A N C E

The fuel pump inlet preignition pressure and temperature were 32.1N/cm 2(46 .5 psia) and 27 8°K (40 .7 °F ), respective ly. T hese fue l pum p in let condi-

tions were w ith in the F - I E ng ine M odel S pecifica tion lim its (sta rt box

requirem ents) as shown in F igure 5-I.

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5-5



5.4 S -IC E N G IN E S HU T DO WNT R A N S IE N T P E R F O RM AN C E

C E C O w as in itia ted by a signa l from the IU at 1 35.1 6 seconds as p lanned.O E C O w as in itia ted by L O X low level sensors and occurred 1 .43 seconds

la ter than the predicted tim e of 1 6 0 .20 seconds tha t w as used in theF ina l F lig h t T ra jecto ry. T h is tim e is w e ll w ith in th e 3 s igm a ra nge fo rO E C O of ± 7 .0 5 seconds. H ow ever, based on the la test pre launch S -IC pro -

pulsion prediction, OECOoccurred only 0.63 second later than predicted.

T hrust decay o f the F - I eng ines was nom ina l. T he to ta l im pu lse from O E C Oto separa tion was 1 0 ,530 ,0 35 N -s (2,36 7 ,247 Ibf-s) wh ich w as w ell w ith inthe 3 sigm a lim its.

5.5 S - IC S T A G E P R O P E L L A N T M A N A G E M E N T

T he S -IC does no t ha ve a n a ctive P ro pellan t U tiliza tio n (P U ) system . M in i-

mum residuals are obtained by attempting to load the mixture ratio ex-

pected to be consumed by the engines plus the predicted unusable residuals.A lso , a sm all additio na l am ou nt o f usable fu e l (fue l bia s) w as loa ded to

m in im ize m axim um residua ls. A n ana lysis of the usable res idua ls experi-

enced during a fligh t is a good m easure o f the perfo rm ance of the passivePU system. S-IC propellant residuals were within expected limits. Usable

L O X residua ls in the tank and suction ducts were approxim ate ly 7 48 kilo-

g ram s (1 6 50 Ibm ) m ore tha n pred ic ted, as co m pared to 2540 kilog ram s (56 0 0

Ibm ) m ore th an p redicted on A S -50 4. A pp ro xim ate ly 1 0 1 2 k ilog ra m s (2230 Ibm )o f usable fue l res idua ls w ere in the fuel tank a t O E C O . T h is w as 31 89

k ilog ram s (7 0 30 Ibm )'less than the fue l b ias o f 420 0 k ilog ram s (9 26 0 Ibm ).

A sum m ary of the propellan ts rem ain ing a t m a jo r even t tim es is presen tedin T able 5-2.

5.6 S -IC P R E S S U R IZ A T IO N S Y S T E M S

5.6 .1 S - IC F uel P ressuriza tion S ystem

T he fuel tank pressurization system performed satisfactorily keeping u llagepressure within acceptable limits during flight. Helium Flow Control Valves(HFCV's) No. 1 through No. 4 opened as planned and HFCV No. 5 was not re-quired.

T he low flow p repressuriza tion system was com m an ded on at -9 7 seco nds.

H igh flow pressuriza tion, accomplished by the onboard pressuriza tion sys-tem , perfo rm ed as expected. H F C V N o. 1 w as com m anded on at -2.6 5 seconds

a nd wa s supp lem ented by th e h igh flow prepressu riza tio n system u ntil um bil-ical disconnect.

F ue l tank u llag e pressure wa s with in the predic ted lim its throughou t fligh t

as show n in F igure 5-4. H F C V 's N o . 2, 3 , and 4 were com m anded open w ith inacceptable lim its during fligh t by the switch selector. Helium bottle p_es-

sure was 2110 N/cm2 (3060 psia) at -2.75 seconds and decayed to 331N/cm z(480 psia ) at O E C O . T o ta l he lium flowra te and heat exchanger perfo rm ancewere as expected.

5-6



T able 5-2. S - IC S tage P ropellant M ass H is tory

PREDICTED LEVELSENSOR

EVENT DATA RECONSTRUCTED

LOX FUEL LOX FUEL LOX FUEL

Ignition k 1,498,856 647,122 NA* 645,577 1,498,137 645,577

Command(lbml (1,423,254)3,304,412) (1,426,660) NA* (1,423,254) (3,302,827)

Holddown kg 1,468,474 638,797 1,464,974 637,054 1,465,078 636,594

ArmRelease (Ibm) ,229,714) (1,404,463)3,237,432) (1,408,307) (3 (3,229,944) (1,403,448)

CECO (Ibkgm) 211,087 98,730 216,081 98,444 216,491 97,976465,367) (217,663) (476,378) (217,033) (477,281) (216,000)

oE co (]bkgm) 1 7 ,S 7 9 1 6 ,0 2g 1 8,347 1 2.87 4 1 8,41 2 1 2,g4438,756) (35,338) (40,448) (28,383) (40,592) (28,537)

Separation kg 15,326 14,946 .... 16,037 ]1,815

(Ibm) (33,787) (32,950) .... (35,356) (26,047)

Zero Thrust (Ib kgm) 15,018 14,650 .... 15,760 11,49733,110) (32,299) .... (34,745) (25,346)

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

* Not available because the LOX was above the level sensors at this time.

Fuel pump inlet pressure was maintained above the required Net Positive

S uction Pressure (NPS P) during flight.

5.6.2 S-IC LOX Pressurization System

The LOX pressurization system performed satisfactorily and met all per-

formance requirements. The ground prepressurization system maintained

ullage pressure within acceptable limits until launch commit. The on-

board pressurization system subsequently maintained ullage pressure

within the GOX Flow Control Valve (GFCV) band during the flight. The

heat exchangers performed as expected.

The prepressurization system was initiated at -71.99 seconds. Ullage

pressure increased until it entered the prepressurization switch band

zone which terminated the flow at -57.69 seconds. The low flow system

was cycled on two additional times at -39.60 and -12.84 seconds. The

high flow system was commanded on at -4.69 seconds and maintained ullage

pressure within acceptable limits until launch commit.

The LOX tank ullage pressure, shown in Figure 5-5, was maintained within

the required limits throughout flight by the GFCV. The maximum GOX flow-

rate (at CECO) was 24.9 kg/s (55.0 Ibm/s).

5-7



HFCV NO. I OPEN _7 HFCV NO. 3 OPEN

H_V NO.2OPEN _7 HFCVNO.4 OPEN

2z .... i I I l..... _ _ i\ 7PRE ICrED MAXIM_ /

i \ - • /.-- \ / \ ...... I .... 30

\ I \I _

20 \

\% - o-AS-505 FLIGHT DATA 28 "_

\ \I /

\ 24

16 \ /

" -- ,I _ ..... 22

"_'LpREDICTED MINIMUM

1 4

_ _7 W 20

0 20 40 60 80 lO0 120 14O 160

_NGE TIME, SECONDS

Figure 5-4. S-IC Fuel Ullage Pressure

The LOX pump inlet pressure met the NPSP requirements throughout flight.

Engine No. 5 LOX suction duct pressure decayed after CECO as shown in

Figure 5-6. The pressure decay rate was 1.38 N/cm2/s (2.0 psi/s) and was

similar to the decay on AS-503 (I.65 N/cm2/s [2.4 psi/s]) and AS-504

(I.38 N/cm2/s [2.0 psi/s]). The cause of this decay is unknown.

5.7 S-IC PNEUMATIC CONTR OL PR ESSUR E SYSTEM

The control pressure system functioned satisfactorily throughout the S-IC

flight.

Sphere pressure was 2130 N/cm2 (3090 psia) at liftoff and remained steady

until CECO when it decreased to 2055 N/cm2 (2980 psia). The decrease was

due to center engine prevalve actuation. There was a further decrease to

1782 N/cm2 (2585 psia) after OECO. Pressure downstream of the regulator

initially read 530 N/cm2 (768 psia) and decreased to 520 N/cm2 (755 psia)

at 160 seconds. R egulator performance was within limits. There were

slight dips in outlet pressure at CECO and OECO due to prevalve actuation.

These dips are to be expected.

The engine prevalves were closed after engine cutoff as required. Engine

No. 5 prevalves closed at approximately 137 seconds. The prevalves for

the other four engines closed at approximately 163.7 seconds.

5-8



l I ///j_ACTIVE PRESSURE/_'j_RELIEF SWITCH 4530 _/._ANDURINGECHANICAL

RELIEFBAND _/'_/'/_LIGHT

............ _16.2 to 17.6 N/cm2 I 40

........7...(23.5 1:o25.6psig

26 ...... ...." l"'-- RELIEFSWITCH

"-, BAND

"'"", "'.. 20.5 to 21.7 N/cm2 35

./_'(29.7 to 31.5 psia,22... ""'"' ].,:.

PREPRESS r_..<_,

_f-S.ITCHAND _>>_._

I - 2s

-- _/7 , ,S-505FLIGHTDATA GFCVBAND-_/

i0 I I I I I _ 15D 20 40 60 80 I00 120 140 ]60

_NGE TIME , S EC ON DS

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

5.8 S-IC PUR GE S YS T EMS

Performance of the S-IC purge systems was satisfactory during the 162-

second flight.

The turbopump LOX seal storage sphere pressure was within the limits of

1903 to 2275 N/cm2 (2760 to 3300 psig) until ignition, and 2275 to 689

N/cm2 (3300 to lO00 psig) from liftoff to cutoff. Regulator outlet pres-

sure remained within the 59 ±7 N/cm2 (85 ±lO psig) limits. Turbopump

LOX seal pressure at the engine interface was within the required limits

of 69 N/cm2 (100 psig) maximum to 21 N/cm2 (30 psig) minimum. The radia-

tion calorimeter purge operated satisfactorily throughout flight.

5.9 P O G O S U P P R E S S IO N S Y S T E M

T he P O G O suppression system perform ed satisfa ctorily prio r to a nd du ring

S -IC fligh t. T he system was in itia lly tu rned on approxim ate ly 26 m inutesp rio r to la un ch to be su re the p reva lves w o u ld fill w ith he lium . R edlinemeasurements indica ted that the four outboard lines filled asscheduled.

T he pressure m easurem ent downstream of the solenoid va lves indicated that

flow was properly established in the system . E leven m inu tes prior tolaunch , the system w as tu rned on aga in and flow was established. T he tem -

perature m easurem ents did not ch ange since the system still con ta ined heliu m

5-9



0 0

0 20 40 60 80 lO0 120 140 160

RANGETIME,SECONDS

Figure 5-_. S-IC Center Engine LOX Suction Line Pressure

from the earlier initiation. The four resistance thermometers performed

as expected during flight. In the outboard lines, the three upper mea-

surements went cold momentarily at liftoff, indicating that the LOX levelshifted on the probes. The probes remained warm throughout flight, in-

dicating helium in the prevalves. Figure 5-7 shows a plot of liquid

level in the prevalve. At cutoff, the increased pressure forced LOX

into the prevalves. The fourth resistance thermometer, at the lip of

the valve cavity, was cold throughout flight as expected.

5-I0



6O -- %_70E C O -P R E VAL VE T O P

_ _l .[ _-_-_-'--

so I l 20PR OB E 1 _-.

i_ _ _o. ___ _ _40 7 .

BE 3_.__ I ,_'/ 15

30 _i" _ SPLASHINGDURINGENGINESTARTUP _

_/_/ LOSS OF CAVITY -- _AFTEROECO -_ 10

20 _ PROBE_ _

• _'

0 _ _/_/_7_ _/_Jv/_ _ _./.PREVALVEOTTO///A////v///.r//M_ / __ _ 0

0 20 40 60 80 100 120 140 160

RANGETI_, SECONDS

Figure 5-7. S-IC Prevalve Liquid Level, Typical Outboard Engine

5-II/5-12



S ECTIO N 6

S -II P R O P ULS IO N

6 .1 S UM M A R Y

The S -II propulsion system performed satisfactorily throughout the flight.

A s sensed by the engines, E ngine S tart Com m and (E S C) occurred at 1 6 3.0 5seconds and O utboard E ngine Cutoff (O E CO ) at 552.6 4 seconds with a burntim e of 389 .59 seconds or 1 .7 0 seconds longer than predicted. T he pre-

dicted propulsion performance parameters used in this section are based on

a revised prelaunch S -II propulsion performance prediction, which was notincorporated in the A S -505 L aunch Vehicle O perational F light T rajectory(dated A pril 1 7 , 1 9 6 9 ). D ue to center engine low frequency perform anceoscillations on the two previous flights, the center engine was shut downearly on A S -50 5 successfully avoiding these oscilla tions. Center E ngineCutoff (CE C O ) occurred at 46 0 .6 1 seconds. T ota l stage thrust, as deter-mined by computer analysis of telemetered propulsion measurements at 6 1seconds after S -II E S C was 0 .35 percent below predicted. T ota l enginepropellant flowrate (excluding pressurization flow) was 0 .43 percent belowpredicted and average specific impulse was 0 .09 percent above predictedat th is tim e slice. A verage E ngine M ixture R atio (E M R ) was 0 .1 8 percentbelo w predicted.

The propellant management system met all performance requirements. T hesystem differed from A S -50 4 by using open-loop control of the engineP ropellant Utilization (P U) va lves. O pen-loop control was utilized onA S -50 3 and is planned for a ll subsequent flights. T he P U valve m ovem entresulted in an actual EMRshift at 488.48 seconds. OECO,initiated bythe L O X low level cutoff sensors, was achieved fo llowing a planned 1 .5-second tim e delay. A sm all engine perform ance decay was noted just priorto cutoff but was less severe than that observed on A S -50 4 due to only

four engines burn ing at cutoff. R esidual propellant remaining in thetanks at O E CO s ignal were 27 89 kilogram s (6 1 50 Ibm ) com pared to a predic-tion of 26 22 kilogram s (57 82 Ibm ).

T he performance of the L O X and L H2 tank pressurization systems was satis-factory. U llage pressure in both tanks was m ore than adequate to m eetengine inlet Net Positive Suction Pressure (NPSP) requirements throughoutmainstage. A s comm anded by the Instrum ent Unit (IU), step pressurizationoccurred at 261.62 seconds for the LOX tank and 461.61 seconds for the

LH2 tank. T he recirculation, engine servicing, pneumatic control andhelium injection systems all performed satisfactorily.

6 - I



6.2 S-If CHILLDOWN AND BUILDUP TRANSIENT PERFORMANCE

The prelaunch servicing operations satisfactorily accomplished the engine

conditioning requirements. T hrust chamber temperatures were within pre-

dicted limits at launch and engine start. The thrust chamber temperatures

ranged between 113 and ]33°K (-256 and -221°F) at -187 seconds, 92 andll3°K (-294 and -256°F) at prelaunch commit and 124 and 145 K (-236 and

-198°F) at engine start. Thrust chamber warmup rates during S-IC boost

agreed closely with those experienced on previous flights.

Both temperature and pressure conditions of the J-2 engine start tanks

were within the required prelaunch and engine start boxes as shown in

Figure 6-I. Start tank temperatures at prelaunch and engine start aver-

aged 13°K (23°F) warmer than on AS-504 as a result of the start tankservicing facility vent line modification. (The vent line flow area was

increased by adding a 3.81 centimeter (I.5 in.) diameter line parallel

to the existing vent line.) Results of this vent line change were highly

satisfactory in that the increased flow area permitted a higher flowrate,

and thus less cooling of the start tank prechill gas passing through theGround Support Equipment (GSE) LH2 heat exchanger. During S-IC boost,

start tank pressureincreaserates due to heatupaveraged3.4 N/cm2/min

(5 psi/min)less than AS-504 results.

S TA RT T AN K T EMPE RA TUR E, O F

-300 -250 -200 -150

1050 I I I

- -15000 ENG NO. l

ZLENGNO.2 r ENGINETARTOX

lO00 Q ENG NO. 3 1 ,r

o ENGNO.4 ! "/--RE LAUN CH BO X

_ ENG NO. 5 _ P/ i1400 ._

950 / .... _j{_11- .

m / J S-II ESC-_-900 -,' | a.A- ._3oo:

#- IVJ

, -PRELAUNCH

850 i n -- _-

_- ' _ 1200

800

750 , , II0080 90 100 llO 120 130 140 150 160 170 180

START TANK TEMPERATURE,°K

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

6 - 2



All engine helium tank pressures were within the required prelaunch and

enginestart 'limitsof 1931 to 2379 N/cm2 (2800to 3450 psia).

During flight, engine No. 5 regulator outlet pressure shifted from 281

to 276 N/cm2 (408to 400 psia) after approximately63 secondsof S-II

engineoperation. Regulatoroperatingrange is 276 ±17.2 N/cm2 (400±25 psia). The engine helium tank pressure also showed a change in decayrate at the same time the regulator shift occurred. Prior to the pressure

regulator shift, the engine No. 5 helium tank pressure decay rate was

3.8 N/cm2/s (5.5 psi/s)comparedto the other enginesdecay rate of about

1.9 N/cm2/s (2.8 psi/s). Subsequentto the shift the helium tank pres-

sure decay rate was 0.57 N/cm2/s (0.83psi/s)which is comparableto the

decay rate of the other engines during the same time period. Even if the

initial decay rate had been sustained throughout S-II burn, the supply

pressure would have been sufficient to meet system demand.

A similar engine helium regulator shift occurred on engine No. 3 during

AS-504 flight. The regulator outlet pressure shifted from 279 to 276

N/cm2 (405 to 400 psia)after approximately43 secondsof engineoperation.

Engine No. 3 helium tank pressure also showed a change in decay rate at

the same time the regulator shift occurred. Prior to the shift the decay

rate was 4.2 N/cm2/s (6.1 psi/s) comparedto about 1.9 N/cm2/s(2.8 psi/s)

for the other engines. Subsequent to the shift, the helium tank pressure

decay rate was 0.76 N/cm2/s(l.l psi/s).

R egulator outlet pressure shifts also occurred on AS-501 and AS-502

flights, but the helium tank pressure decay rate did not change at the

same time. The regulator shifts were not experienced on AS-503. The

probable cause of these minor regulator shifts and changes in engine

helium tank pressure decay rates is internal regulator leakage.

The LOX and LH2 recirculation systems used to chill the feed ducts, turbo-

pumps, and other engine components performed satisfactorily during pre-

launch 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 ESC were 6.2 to 7.3°K (ll.2 to

13.2°F ) subcooled, which is well below the 1.7°K (3°F ) subcooling require-m e n t .

Prepressurization of the propellant tanks was satisfactorily accomplished.

Ullagepressuresat S-II ESC were 26.9N/cm2 (39 psia) for LOX and 19.3

N/cm2 (28 psia) for LH2.

S-II ESC was received at 163.05 seconds and the Start Tank Discharge Valve

(STDV) solenoid activation signal occurred l.O second later. The engine

thrust buildup was satisfactory and was within the required thrust build-

up envelope. The stage thrust reached mainstage level at 166.30 seconds.

E ngine thrust levels were between 854,059 and 898,541 N ewtons ( 192,000and 202,000 Ibf) prior to " High EMR Select" command at 168.50 seconds.

6-3



L H2 P UM P IN LE T P RE SS UR E, psia28 32 36 40 44

/ -415

SATURATIONCURVE--.F/_/° 24 _ /J -417

/ -419_ 22 f_

/-PREDICTED ESC

' r_-_FLIGHT DATA _

ALL ENGINES -423

20

-425

1 9

18 20 22 24 26 28 30 32

LH2 PUMPINLET PRESSURE, N/cm2

L O X P UM P IN LE T P RE S SUR E, psia

32 36 40 44 48

101 ST/AR_SAT URATION -2782809 ,-- CURVE /

L J -

0

N R-284_

_- ENVELOPE/97

_ -286_

LU -288_- 95 _

w / PREDICTEDESC -290 _-w"-J / _ .=J

/ -292

_ 93 / rFLIGHT DATA

_ J( /ALL ENGINES -294 _x

o 91 i ._j -296-298

8 9

21 23 25 27 29 31 33 35

LOX PUMPINLET PRESSURE,N/cm2

F igure 6 -2. S -II E ngine P um p In le t S ta rt R equ irem ents

6-4



6.3 S-ll MAINSTAGE PERFORMANCE

Stage performance during both the high and low EMR portions of the flight

was very close to predicted as shown in Figure 6-3. At the ESC +61-secondtime slice, total vehicle thrust was 5,157,611 N ewtons ( I,159,477 Ibf)

which is only 18,064 Newtons (4061 Ibf) or 0.35 percent below the pre-

flight prediction. Average enginespecificimpulsewas 4165.2 N-s/kg(424.7Ibf-s/Ibm)or 0.09 percentabove the predictedlevel. Propellant

flowrateto the engines (excludingpressurizationflow) was 1238.3kg/s

(2729.9Ibm/s)which was 0.43 percentbelow prediction,and the averageE MR was 5.56 or 0.18 percent below preflight prediction.

At ESC +297.56 seconds, the center engine was shut down in order to ore-

vent buildup of the low frequency oscillations that were observed onAS-503 and AS-504. This action reduced total vehicle thrust by 1,044,060

Newtons (234,714 Ibf) to a level of 4,103,720 Newtons (922,553 Ibf). Ofthis total, 1,024,274 N ewtons (230,266 Ibf) were directly due to CEC O and

the remaining 19,786 Newtons (4448 Ibf) resulted from the effect of fuel

step pressurization and loss of acceleration head.

The PU system was operated in the open-loop control mode for the AS-505

flight. At approximately 325 seconds after ESC, engine thrust chamber

pressures reacted to the PU control valve step from the high to low EMR

position. The action further reduced total vehicle thrust to 3,090,002

Newtons (694,660 Ibf) at ESC +350 seconds. A change in stage thrust of

1,013,576 Newtons (227,861 Ibf) is indicated between high (5.47) and low

(4.27) EMR operation. Unlike previous flights, the deviation of actual

from predicted performance did not increase at the lower mixture ratio

levels. Vehicle thrust and propellant flowrate deviations at ESC +388

secondswere -ll,161Newtons (-2509Ibf) and 1.8 kg/s (3.9 Ibm/s),respec-

t i v e l y .

Individual J-2 engine data, excluding the effects of pressurization flow-

rate, are presented in Table 6-I for the ESC +61-second time slice. With

the exception of engine No. 5, very good correlation between predictionand flight was indicated by the small deviations. Flight data reconstruc-

tion precedures were directed toward matching the engine and stage accept-

ance specific: impulse values while maintaining the engine flow and pump

speed data as a baseline.

Examination of engine No. 5 data indicated that the low performance level

resulted from a large increase in Gas Generator (GG) LOX bootstrap line

hydraulic resistance. This lower engine power level was maintained

throughout the flight. During vehicle acceptance testing, this engineexhibited two short intervals of reduced performance of 20,017 and 14,679

Newtons (4500 and 3300 Ibf) thrust. Following the static test operations,

a complete inspection was performed of the GG injector, control valve and

bootstrap line. No contamination, restrictions, or out-of-toleranceconditi ons wer e de tected.

6 - 5



9-9

STAGE

SPECIFIC

IMPULSE,

IO

N-s/

g

STAGE

THRUST,

10

N

_

i

0

omm

_

N

0

!

m

_

N

N

e

_

°N

N o

'

o

o

o

o

_

g

_

N

_

N

o

o

o

S

S

I

Ibf-

s/

_

c

STAGE

AVENGE

ENGINE

MIXTURE

_TIO,

LOX/

UEL

STAGE

TOTAL

FLOWRATE,

kg/

o

N-

_

_

=_

._

._

_

_

_

_

_

o

o

o

o

_

_

_

_

_

o

o

o

o

o

o

o

O

0

o

o

_

__ om

mmm



T able 6 -I. S -II E ngine P erform ance Deviations (E S C +6 1 S econds)

PERCENT AVERAGEPERCENT

PARAMETER ENGINE PREDICTED RECONSTRUCTION DEVIATION DEVIATIONANALYSIS

FR OM PR EDICTED FR OM PR EDICTED

1 1,036,316 (232,973) 1,036,098 (232,924) -0.02

Thrust, 2 1,032,704(232,161) 1,027,855(231,071) -0.47

3 1,028,002 (231,104) 1,031,276(231,840) 0.32 -0.35

Newtons(Ibf) 4 1,036,004(232,903) 1,038,553(233,476) 0.255 1,042,650 (234,397) 1,023,829(230,166) -I.81

I 4161.9 (424.4) 4169.8 (425.2) 0.19

Specific 2 4157.0 (423.9) 4162.9 (424.5) 0.14Impulse, 3 4154.1 (423.6) 4152.1 (423.4) -0.05 0.09

N-s/kg(Ibf-s/Ibm) 4 4161.9 (424.4) 4168.8 (425.1) 0.165 4172.7 (425.5) 4172.7 (425.5) 0

I 249.0 (548.9) 248.5 (547.8) -0.20Flowrate, 2 248.4 (547.7) 246.9 (544.3) -0.62

3 247.5 (545.6) 248.4 (547.6) 0.37 -0.43

kg/s (Ibm/s) 4 248.9 (548.7) 249.2 (549.3) 0.115 249.9 (550.9) 245.3 (540.9) -1.82

I 5.58 5.53 -0.90Mixtureatio, 2 5.58 5.58 0

3 5.63 5.63 0 -0.18LOX/Fuel 4 5.55 5.55 0

5 5.51 5.51 O

Actual flight data are presented in Table 6-I 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. In

comparison to the vehicle acceptance test, the adjusted data showed

engines No. 2 and 5 to be l.O and 1.86 percent low in thrust, respectively.

The low frequency oscillations which occurred on AS -503 and AS -504 didnot occur on this flight. The oscillation problem appeared to be associ-

ated with inflight LOX liquid levels. The LOX level history for all S-II

stage flights is shown in Figure 6-4. Early cutoff of the center engine

on AS-505 precluded any oscillation buildup. Subsequent to CECO no

adverse structural response characteristics were evident (for a detailed

discussion refer to Section 9, paragraph 9.2.3). The flight results verify

that early cutoff of the center engine successfully avoided the low fre-

quen cy osci lla tio n p roblem .

6 .4 S -II S HUT DO WNTR AN SIE N T P ER FO RM AN CE

E ngine shutdown sequence was initiated by the stage L O X low level sensors.

T he O E CO signal was delayed 1 .5 seconds after the low level sensor dryindications by tim ers in the L O X depletion cutoff system. T his resulted

in engine performance decay prior to receipt of the cutoff signal, similarto that experienced during A S -50 4 flight. Due to early CE CO h owever, theprecutoff decay was greatly reduced. O nly engine N o. 1 exhibited a signif-icant thrust chamber pressure decay prior to cutoff, decreasing approxi-mately 79.3 N/cm2 (115 psi) in the final 1.5 seconds. The decay of thrust

6 - 7



45.75 STARTFENGINE 1804

VOSCILLATION PERIOD

45.50 _ S-II EMRSHIFT 1794 i

E45,00 \\ _ 1774_

_4,75_ _ , _ 1764 _\ AS502

44.50 AS-505 1754

• _'>f_ AS-504

44.00 1734

AS-501

43.75 _7 _ I

420 440 460 480 500 520 540 560 5801724

R_GE TIME, SECONDS

Figure 6-4. S-II Inflight LOX Level History

chamber pressures of the other outboard engines was approximately 13,8

N/cm2 (20 psi). The order of outboard engine thrust chamber pressure

decay was identical for the AS-504 and AS-505 flights (engines No. l, 2,

3 and 4). One second before cutoff, with all outboard engines operating

at low mixture ratio, total stage thrust was approximately 3,062,102

Newtons (688,388 Ibf) with an average specific impulse of 4221.8 N-s/kg(430.5 Ibf-s/Ibm).

6-8



A t O E CO (552.6 4 seconds) the tota l vehicle thrust was down to 2,856 ,0 6 1N ewtons (6 42_.0 6 8 Ibf). Vehicle thrust dropped to 5 percent of th is levelwithin 0 .6 6 second. T he stage cutoff im pulse through the 5 percentthrust level was estim ated to be 56 3,1 45 N -s (1 26 ,6 0 0 Ibf-s). G uidance

data indicates the total impulse from OECOto S-II/S-IVB separation at553.50 seconds to be 57 8,7 1 4 N -s (1 30 ,1 0 0 Ibf-s) com pared to a predicted

value of 6 47 ,,7 50 N -s (1 45,6 20 Ibf-s) for th is tim e period. N o unusualfeatureswere apparentin the center engine thrustdecay data followingCECO. The 5 percent thrust level was reached approximately 0.3 secondafter cutoff.

Based on the latest propulsion performance prediction, burn time was 1.70seconds longer than expected. A comparison of flight data with the

L aunch Vehicle O perational F light T rajectory ( dated A pril 17, 1969), which

was based on a previouspropulsionprediction,indicatesa differentburntime deviation.

6.5 S-II STAGE PROPELLANT MANAGEMENT

The propellant management system performed satisfactorily during the pro-

pellant loading operation and during flight. The S-II stage employed an

open-loop system utilizing fixed, open-loop commands from the IU rather

than feedback signals from the tank mass sensing probes. (Open-loop

operation was also used on AS-503 and is planned for use on all subse-

quent vehicles).

The facility Propellant T anking C ontrol S ystem (PT CS ) and the propellant

management system successfully accomplished S-II loading and replenish-

ment. During the prelaunch countdown, all propellant management sub-

systems operated properly with no problems noted. Propellant fill and

drain valve closure times were satisfactory (8.97 seconds for the LOX

valve and 18.64 seconds for the LH2 valve).

DuringCDDT, splashingof the LH2 overfillshutoff(liquidlevel)sensoroccurred. After the LH2 tank was filled and the replenish mode was

established during countdown, the GSE " revert" interlock for this sensorwas deactivated.

Open-loop PU system operation commenced when " High EMR Select" was com-

manded at ESC +5.45 seconds as planned. The PU valves then moved to the

high EMR position, providing a nominal EMR of 5.50 for the first phase

of Programed Mixture R atio (PMR ). No propellant management system

anomalies resulted from CECO. At ESC +323.49 seconds, the low EMR command

was initiated, driving the PU valves against the low EMR stop. This pro-

vided an average EMR of 4.31 to l (predicted 4.32 to l) for the remaining

low mixture ratio portion of the flight.

The open-loop PU control system responded as expected during flight and

no instabilities were noted. The open-loop PU error at OECO was

6 - 9



approximately -31 .7 kilogram s (-7 0 Ibm ) L H2 versus a 3 sigm a toleranceof ±1 1 34 kilogram s (±250 0 Ibm ).

Based on point level sensor data, propellant residuals (mass in tanks andsum ps) at O E CO were 81 6 kilogram s (1 80 0 Ibm ) L O X, and 1 9 7 3 kilogram s

(4350 Ibm ) L H2, versus the predicted 6 56 k ilogram s (1 447 Ibm ) L O X, and1 9 6 6 kilogram s (4335 Ibm ) L H 2. A n updated analysis using A S -50 4 L O × deple-tion data indicated a higher L O X residual would result on A S -50 5. Correc-tions for CE CO a nd E M R differences resulted in a revised L O × predictedresidual o f 7 80 kilogram s (1 7 21 Ibm ). T able 6 -2 presents a com parison ofpropellant m asses as measured by the P U probes, engine flowmeters andpoint level sensors. T he best estimate propellant m ass is based on inte-

gration of flowmeter data utilizing the propellant residuals determinedfrom point level sensor data at O E CO . B est estimates of propellant m assloaded are 37 2,71 7 kilograms (821 ,7 0 0 Ibm ) L O X, and 71 ,80 8 kilogram s

(1 58,31 0 Ibm ) L H2. T hese m ass values are 0 .1 4 percent less than predictedfor L O X and 0 .20 percent more than predicted for L H2.

T able 6 -2. S -II P ropellant M ass History

EVENT PUSYSTEM FLOWMETERNALYSIS POINT

" SENSOR

RANGETIME UNITS PREDICTED ANALYSIS (BESTESTIMATE) ANALYSIS

LOX LH2 LOX LH2 LOX LH2 LOX LH2

Ground kg 373,249 71,668 373,218 71,599 372,7]7 71,808 372,866 72,|97

Ignition (]bm) (822,874) (158,000) (822,805) (157,848) (821,700) (158,310) (822,028) (159,168)

S-II ESC kg 373,249 7],668 373,004 71,548 372,717 71,808 372,866 72,197

{163.05sec) (Ibm) (822,874) (158,000) (822,332) (157,737) (821,700) (158,310) (822,028) (159,168)

High EMR Select k

(ibmg) 370,463 70,9]8 369,852 7],078 370,166 71,]47 369,367 72,082168.50 sec) (816,731) (156,348) (815,384) (156,700) (816,076) (156,852) (814,314) (158,914)

PU Va lve S te p

(lbkgm) 38,290 ]0,74] 49,936 ]1,007 41,093 11,897 42,5]2 11,495486.54

sec) (84,415) (23,679) (II0,089) (24,267) (90,594) (26,229) (93,723) (25,343)

S-IIOECO k

(ibmg) 656 1966 1433 1694 816 1973 816 1973552.64 sec) (1447) (4335) (3160) (3735) (1800) (4350) (1800) (4350)

S-ll Residual k

(Ibm9 542 1916 1305 1651 689 1930 689 1930fter Thrust Deca) (I194) (4224) (2878) (3639) (1518) (4254) (1518) (4254)

NOTE: This table does not include propellant trapped external to the tanks and LOX sump.

6.6 S -II P R E S S UR IZA T IO N S Y S T E M S

6 .6 .1 S -II Fuel P ressurization S ystem

L H2 tank ullage pressure, actual and predicted, is presented in F igure 6 -5for autosequence, S -IC boost and S -II boost. T he L H2 tank vent va lveswere closed at -96 seconds and the ullage was pressurized to 24.6 N/cm2(35.7 psia) in approxim ately 25 seconds. O ne m akeup cycle was requiredat -38 seconds as a result o f therm al pressure decay. Venting occurredduring S -lC boost as anticipated. T wo venting cycles were indicated onvent valve N o. 1 between 6 3 and 88 seconds. T here was no indica tion that

vent valve N o. 2 opened. Differentia l pressure across the vent va lve was

6-10



kept below the low-mode upper limit of 20.3 N/cm2 (29.5 psid). Ullagepressure at S-II engine start was 19.3 N/cm2 (28 psia) meeting the mini-mumengine start requirement of 18.6 N/cm2 (27 psia). The LH2 tank valveswere switched to the high vent mode imm ediately prior to S -II eng ine start.

LH2 tank ullage pressure was maintained within the regulator range of19.7 to 20.7 N/cm2 (28.5 to 30 psia) during burn until the LH2 tank pres-sure regulator was stepped open at 461 .61 seconds. Ullage pressureincreased to 22.1N/cm2 (32 psia). The LH2 vent valves started ventingat 483 seconds and continued venting throughout the remainder of the S -II

flight. Ullage pressure remained within the high-mode vent range of 21to 22.7 N/cm2 (30.5 to 33 psia).

Figure 6-6 shows LH2 total inlet pressure, temperature and NPSP. The

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 L.OXPressurization System

LOX tank ullage pressure, actual and predicted, is presented in Figure£-7 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

closedat -185.4secondsand the LOX tankwas prepressurizedto the

pressureswitch settingof 26.6 N/cm2 (38.6 psia) in approximately50

seconds. No pressure makeup cycles were required. However, a slight

pressure decay occurred, which was followed by the slight pressure in-

crease caused by L H2 tank prepressurization. Ullage pressure was 26.9N/cm2 (39 psia) at enginestart.

The LOX regulator remained at its minimum position until 245 seconds

because the ullage pressure was above the acceptable regulator range of24.8 to 26.5 N/cmZ (36 to 38.5 psia). A slight decreasein ullagepres-

sure prior to LOX regulator step pressurization indicated normal per-

formance of the L OX regulator. L OX step pressurization (261.62 seconds)

caused the usual characteristicsurge in ullagepressurefollowedby a

slower increaseuntil EMR shift. LOX tank ullagepressurereachedamaximumof 28.2 N/cm2 (40.9psia)beforethe characteristicdecaywhich

followsEMR shift. Ullage pressurewas 25 N/cm2 (36.3psia) at OECO.

LOX pump total inlet pressure, temperature and NPSP are presented in

Figure 6-8. The NPSP supplied exceeded the requirement throughout the

S-II boost phase. The total magnitude of LOX liquid stratification was

slightly greater than predicted. The 1.5-second time delay in the LOX

low levelcutoff circuitused for AS-504and AS-505makes it very diffi-cult to predict an accurate cutoff temperature.

6.7 S-II PNEUMATIC CONTROl. PRESSUR E 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

6 - 1 1



_7 V E N T V A L V E N O . 1 O PE N (F IR S T )

ACTUAL _7 VENTVALVENO. I CLOSESECOND)-----PREDICTEO S_7S-IIESC

S-IICECO,460.61 45

30 _ LN2STEPPRESSURIZATION,61.61ENT VALVES OPEN

_' S-II OECO,552.64 40

% _7 VENT VALVES CLOSED, (POST OECO) ._

25

REDL NE -- 3O20 "

MINIMUMTART 25

REQUIREMENT

15

20

1C _7_ _7 _7_7 _ 15

-200 -lO0 0 I00 200 300 400 500 600


Figure 6-5. S-II Fuel Tank Ullage Pressure

the operatingbandof 476 to 527"N/cm2(690 to 765 psia)except duringvalve actuations which follow S-II ESC, CECO and OECO events. The makeup

period for the regulator outlet pressure to return to its operating bandafter 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.

Pressuredecreasedfrom 2065 to 2062 N/cm2 (2995 to 2990 psia) during

this period. Main receiverpressurewas 1817N/cm2 (2635psia) followingthe final valve actuation at OECO.

6 .8 S -II HE L IUM IN JE CT IO N S Y S T E M

T he performance of the helium injection system was satisfactory. R equire-ments were met and parameters were in good agreement with predictions.The supply bottle was pressurized to 2068 N/cm2 (3000 psia) prior to lift-

off and by ESCwas 448 N/cm2 (650 psia). Helium injection system averageto tal flowrate during supply bottle blowdown (-30 to 1 6 1 .75 seconds) was2.0 1 S C M M (7 0 .7 S C F M ).

6-12



S-IIESC _LH 2 STEPPRESSURIZATION,461.61

S-IICECO, 460.6] _S-II OECO

25 36_ --ACTUAL_E

o .... PREDICTED < _

_z 23 34 _W _

_ ...... 30 _

19 28

2 3_z -419u_0 0

22 -420u]

_ 21 _--_--- _ -422 _('J:E _-

_ -423_

2 O

1 2

1 6

u 12"_ 8

= =3

_- 6 z

MINIMUMNPSP REQUIREMENT 6

42

0 50 IO0 150 200 250 300 350 400

T IM E F R O M E S C, S E CO NDS

163 263 363 463 563

R ANGE TIME, SECONDS

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

6-13



ACTUAL _ S-II ESC.... P R E DICT E D

_7 L O X S TE P P R E S S U R IZ A T IO N

X_ S-II CECO -453 O

_ S-II OECO

-- ..... t ..J_C_ __o5 _

-35

_ LAUNCH CL _"REDLINE -_"_ u_z_c_

MNMUM

, __ STARTEQUIREMENT --30

× -25×0

1 5-20

i_o_ V '_' :7 ,_, __-200 -I00 0 I00 200 300 400 500 600

R AN GE T IM E, S ECO NDS

F ig u re 6 -7 . S - I I L O X T a nk U llag e P ressu re





S E CT IO N 7

S -IVB P R O PULS IO N

7 .1 S UM M A R Y

The J-2 engine operated satisfactorily throughout the operational phaseof first and second burns. S hutdowns for both burns were also norm al.

T he engine performance during first burn, as determined from standardaltitude reconstruction analysis, was 0 .13 percent less than predicted

for thrust and 0 .26 percent greater than predicted for specific impulse.T he first burn duration was 1 46 .95 seconds from S tart T ank Discharge

Valve (S T DV) open. T his duration was 1 .54 seconds longer than predicted.

E ngine C utoff (E CO ) was in itia ted by a velocity cutoff com m and from theL aunch Vehicle Digita l Computer (L VDC).

T he Continuous Vent S ystem (CVS ) adequately regula ted L H 2 tank ullagepressure at 13.4 N/cm2 (19.5 psia) during earth parking orbit. TheOxygen/Hydrogen (02/H 2) burner satisfactorily repressurized the LH2 tankfor restart. R epressurization of the L O X tank was not required.

E ngine restart conditions were with in specified limits. R estart at fu ll

open P ropellant Utilization (P U) valve position was successful and therewere no indications of overtem peratures in the Gas Generator (GG).

S econd burn duration was 343.0 6 seconds from S T DV open which was 0 .65

second shorter than predicted. E ngine performance during second burn,as determined from the standard altitude reconstruction analysis, was

0 .25 percent less than predicted for thrust and 0 .30 percent greater thanpredicted for specific im pulse. E CO was initia ted by a L VDC velocitycutoff command.

S ubsequent to second burn, the propellant lead experiment was success-fu lly accomplished and the stage propellant tanks and pneumatic systemswere satisfactorily safed. T he velocity change resulting from the

experim ent, CVS , the L O X dum p, and A uxilia ry P ropulsion S ystem (A P S ) firingscaused the stage to enter a solar orbit as p lanned.

A helium leak in the A P S m odule N o. 1 was noted at 23,40 0 seconds

(0 6 :30 :0 0 ). T he leak persisted until loss of data at 39 ,240 seconds(1 0 :54:00 ); however, system performance was with in operational limits.

7 -I



7 .2 S -IVB C H IL L DO WN A N D B U IL DU P T R A N S IE N T P E R F O RM AN C E F O R F IR S T B U R N

T he propellant recircula tion systems performed satisfactorily and met

sta rt and run box requ irem ents fo r fue l and L O X as show n in F igure 7 -I.T he th rust cham ber tem pera tu re at launch was well below the m axim um

allow able red lin e lim it o f 1 7 2°K (-1 50 °F ). A t S - IV B firs t burn E ng ine

S ta rt C om m an d (E S C ), th e tem pera tu re w as 1 59 .4°K (-1 7 3°F ), w h ich is

within the requirement of 166 ±27.5°K (-160.9 ±49.5°F).

T he ch illdow n and loading o f the eng ine G aseous H ydrogen (G H 2) sta rt

sphere and pneum atic contro l sphere prio r to lifto ff were satisfactory.A t first E S C the start tank conditions were with in the requ ired S -IVB

region of 896.3 ±68.9 N/cm2 (1300 ±I00 psia) and 133.2 ±44.4°K (-220±80°F) for initial start. The discharge was completed and the refillin itia ted at first burn E S C + 4.40 seconds. T he refill was sa tisfacto ry

and in go od agreem ent with the acceptan ce test.

T he enqine contro l bottle pressure and tem perature a t lifto ff were 20 82N/cm2 _3020 psia) and 178°K (-140°F), respectively.

L O X and L H 2 system s ch illdowns, which were contin uous fro m before lifto ffuntil ju st p rio r to S - IV B firs t bu rn E S C , w ere sa tisfac to ry . A t E S C th e

L O X pum p in le t tem p era tu re w as 9 1 .3°K (-29 5 .5°F ) and the L H 2 p um p in le t

tem p erature wa s 21 .4°K (-421 .5°F ).

T he first burn sta rt transient was satisfactory. T he thrust buildup was

with in the lim its.set by th e eng in e m anufacturer. F aster thrust bu ildup

to the 9 0 percen t leve l as com pared to the accep tance test resu lts was

observed on th is fligh t. T h is bu ildup was sim ila r to the th rust bu ildupsobserved on prev ious fligh ts . T he P U va lve w as in proper nu ll position

prio r to firs t s ta rt. T h e to ta l im pu lse from S T D V to S T D V + 2.5 secondswas 832,9 43 N -s (1 87 ,253 Ibf-s) fo r first s ta rt. T h is w as greater than

the va lue of 6 44,9 9 2 N -s (1 46 ,0 0 0 Ibf-s) obta ined during the sam e in terva l

for the acceptance test.

A lthough the fuel in jection tem perature m easurem ent behaved in an erratic

m anner, the first burn fuel lead appeared to fo llow predictions. R ela ted

measurements and subsequent performance indicated that satisfactory condi-tions were provided.

7 .3 S -IVB M AIN S T A G E P E R F O RM AN C EF O R F IR S T B U R N

S -IV B stage propu lsion system perform ance is eva luated using propu lsionreconstruction ana lysis. T his ana lysis u tilizes te lemetered eng ine and

stage data to com pute longitudinal thrust, specific impulse, and stage

m ass flowrate. T he propu ls ion reconstru ction ana lysis showed that the

stage perform ance during mainstage operation was satisfactory. A com-

parison of predicted and actual performance of thrust, tota l flowrate,specific im pulse, and m ixtu re ra tio versus tim e is shown in F igure 7 -2.

T able 7 -I sho ws th e specific im pulse , flowrates and m ixture ra tio

7- 2



F UE L PUMP IN LE T PR ES SUR E, psia

16 20 24 2S 32 36 40 44

25 I I , I -415

T IM E F RO M F IR S TE SC ,ITEM SECONDS

24 I O _,_

2 3 (STDV) _._ ! -417

4 10o °5 ECO _._

1 I -421 _

!4 ENGINE START LIMITS _,'_ -423

20 I

-425

19

I0 12 14 16 IB 20 22 24 26 28 30 32

FUEL PUMP INLET PRESSURE. N/cm2

LOX PUMP INLET PRESSU_, psia

28 32 36 40 44 48 52

100 ! i .4 _ -280

TIME FROM FIRST ESC, /(TEM SECONDS /

98 2 3(STDV) ._ -2_3 50

4 lOO

5 _50 I

l_ , -22

342151 0 0 -296

-ENGINE START LIMITS --

/ l ....... J90

-300

8818 20 22 24 26 28 30 32 34 36 J8

LOX PUMP INLET PRESSURE. N/cm2

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

7 - 3



ACTUAL _7 FIRSTESCPREDICTEDBAND _/'FIRST ECO

1 . 0 04 -

z 0.22.Q

o 0.21 ,.o

0.90 - 0.20

c_ m-r

0.19

080 O18

4300

_ _ 435_m- - 4304200 _E

_ 425 "_

_" 420 u_"-4 1 O C

:z 230 I

"_ 500E

_3___,._ _ .....

205 450

_..1-...I LL

I-.- ..j0

_- 180 4000

5.5 _-

_ 5.0

t._ Z2_

zo _x 4.5 0 20 40 60 80 lO0 120 140 160

T IM E FROM STDV +2 .5 SECONDS

550 600 650 700

RA NGE TIM E, SECON DS

F igure 7 -2. S -IV B S teady S ta te P erfo rm ance - F irst B urn

7-4



Table 7-I. S-IVB Steady State Performance - First Burn (ES C +140-Second

Time Slice at Standard Altitude Conditions)

PA R AME T E R PR EDIC T ED R EC O N S T RUC TIO N F LIGHT PER C E N TEVIATIONDEVIATION FROM PREDICTED

ThrustN 9,129,304 9,117,294 -12,010 -0.]3

(Ibf) (205,235) (204,965) (-270)

Specific Impulse

N-s/kg 4201 4212 l] 0.26

(Ibf-s/Ibm) (428.4) (429.5) (l.l)

LOX Flowrate

kg/s 180.91 180.18 -0.73 -0.40

(Ibm/s) (398.83) (397.24) (-I.59)

Fuel Flowrate

kg/s 36.38 36.30 -0.08 -0.24

(Ibm/s) (80.21) (80.02) (-0.19)

Engine MixtureR a t i o

LOX/Fuel 4.972 4.964 -0.008 -0.16

deviations from the predicted at the ESC +140-second time slice when the

engine performance characteristics stabilized. This time slice perform-

ance is the standardized altitude performance which is comparable to

engine tests. The 140-second time slice performance for first burn thrust

was 0.13 percent lower than predicted. Specific impulse performance for

first burn was 0.26 percent higher than predicted.

First burn duration was 146.95 seconds from STDV open, which was 1.54

seconds longer than predicted burn time.

Instrumentation installed to monitor Augmented Spark Igniter (ASI) system

performance responded as expected. Both LOX and LH 2 supply line tempera-

tures chilled to expected levels during both burns and did not indicate

any abnormal conditions.

The helium control system for the J-2 engine performed satisfactorily

during first burn mainstage operation. Since the engine bottle was con-

nected w ith the sta ge am bient repressuriza tion bottles there was little

pressure decay. H elium usage was estim ated from flowra tes during engine

opera tion . A pproxim ate ly 0 .1 54 kilog ram (0 .34 Ibm ) was consum ed duringfirst burn.

7 .4 S -IV B S H U T DO W NT R A N SIE N T P E R F O RM A N CE F O R F IR S T B UR N

T h e E C O transient was satis factory and agreed clo sely w ith th e acceptan cetest and predictions. T he to ta l cuto ff im pu lse to zero percen t o f ra ted

th rust w as 20 3,37 3 N -s (45 ,7 20 Ibf-s). C u to ff occurred with the P U va lvein the nu ll pos ition .

7-5



W hen cuto ff im pulse was adju sted for anticipa ted M ain O xidizer V a lve

(M O V ) tem pera tu re and com p ared w ith the log boo k va lues a t nu ll P U va lve

position and 255°K (O°F) MOVactuator temperature, the flight value wasnear the log book va lue.

7 .5 S -IVB P A R KIN G O RB IT C O AS T P H A S E C O NDIT IO N IN G

T he L H2 CVS performed satisfactorily,main ta in ing the fuel tank u llagepressure at an average level of 13.4 N/cm2 (19.5 psia).

The continuous vent regulator was activated at 762.95 seconds. Regula-tion con tinued with the exp ected opera tion of the m ain popp et periodi-

cally open ing , cycling , and reseating. Continuous venting was term inated

at 86 7 1 .42 seconds. T he C VS perfo rm ance is show n in F igure 7 -3.

Calcula tions based on estimated tem peratures indicated that the m ass

ven ted during park ing orbit w as 1 0 1 4 k ilo g ram s (2236 Ibm ) a nd th a t th e

bo ilo ff m ass w as 1 0 9 2 k ilo g ram s (240 7 Ibm ).

7 .6 S - IV B C H IL L D O W N A N D R E S T A R T F O R S E C O N D B U R N

Propellant tank repressurization was satisfactorily accomplished by the

02/H2 burner. Helium heater "ON" command was initiated at 8671.2 seconds.L O X ta nk u llage pressure at he lium heater "O N " com m an d was approxim ately27.1 N/cm2 (39.3 psia); therefore, repressurization of the LOX tank was

not required. T he L H2 repressurization contro l va lves were opened at

he lium heater "O N " +6 .1 seconds. T he fue l tank was repressurized from13.2 to 20.9 N/cm2 (19.2 to 30.3 psia) in 182 seconds which yielded a

ramp rate of 2.48 N/cm2/min (3.59 psi/min) as shown in Figure 7-4. Therew ere 1 2.7 kilog ram s (28.0 Ibm ) o f co ld helium used from the co ld he liumspheres during repressu riza tion . T he burner continued to opera te for a

tota l of 460 seconds and provided nom ina l propellan t settling forces.

The performance of the 02/H 2 burner was satisfactory as shown in Figure7-5.

T he S -IVB stage provided adequate condition ing of propellants for eng ine

restart. The engine start sphere was recharged properly and maintainedsufficien t pressure during co ast. T he eng in e con tro l sphere gas usag e

was as predicted during the first burn ; the am bient he lium spheres re-charged the control sphere to a nominal level adequate for a proper re-start.

T he propellant recirculation systems performed satisfactorily and met

sta rt a nd run box requ irem ents fo r fue l and L O X as show n in F igure 7 -6 .T h e L H2 pu m p in le t tem pera tu re a t seco nd bu rn E S C w as 23.9 °K (-41 6 .6 °F ).A t S -IVB second burn E S C the L O X pum p in le t tem pera tu re w as 9 1 .1 °K

(-29 5.7 °F ). S econd burn fue l lead genera lly fo llow ed the predictedpattern and resu lted in satisfactory conditions as indica ted by the

thrust cham ber temperatures and the associa ted fuel in jector tempera-

tu res. T he start tank perfo rm ed satis factorily during the second burn

blowdown and recharge sequence.

7- 6



_7 FIRST ECO

20E _

Z _

_ 15 _

I0 m

N 5 NN N

_ 0 0 _

50 0

10 0

400

m 75 _

= 300

_ 50 _

o 200 ,-_

I00 PREDICTED_ "_ 25

0 0

0.0003

0

<_ 0.0002 L

o.oool> 0.0000

0 500 I000 1500 2000 2500 3000 3500 4000 4500 5000

RANGE T IM E, SECONDS

! _7 I I I

0:00:00 0:30:00 1:00:00 1:30:00

R A N GE T IM E , H OU RS :M IN UT ES :S E CO N DS

Figure 7-3. S-IVB CVS Performance - Coast Phase (Sheet 1 of 2)

7 - 7



_T S T A R T O F T 6 _T S E C O N DE S C20

E

U

z

m 20_

_ mm ]0 _

m ]0 __ N

N 5 ---- NN 00 Zz

> 0 0 _

50 0

100

z 400

_2 _2m 75m300 =

< 50200

o jPREDICTED

100 , 25 _

0 0

0.0003

z

0 . 0 0 0 2

o.oooo r--50 0 0 5 50 0 6 0 0 0 6 5 0 0 7 0 0 0 7 50 0 80 0 0 85 0 0 9 0 0 0 9 50 0 I0 ,0 0 0


I , ,57 ,1:30:00 2:00:00 2:30:00 3:00:00

R A N GE T IM E , H O UR S :M IN UT E S:S E CO N DS

F igu re 7 -3. S -IV B C V S P erfo rm an ce - C oas t P hase (S heet 2 o f 2)

7-8





_7 HE L IUM HE A T E R O N

_7 T E R M IN A TIO N O F L H 2 T A NK R E P R E S SS _7 HE LIUM HE AT ER O FF

0.I0

0.20

0.08

-_ DATAAP O.]5 _.JC3

0.06 vf

-J 0.I0J_

0.04 o

i . = 4

JL_J

0.02 0.05m

0.0 0.002 0 0

4 0

150 4_z .c}

_ _r_ 20 _

I.u e_z

= Nff =c_

o Lo-1O0 1O0 200 300 400 500

T IM E F R O M HE L IUM HE A TE R O N, S ECO NDS

, _ i _, _ , _ i2:22:50 2:26:10 2:29:30 2:32:50

R A NG E T IM E , HO UR S :M IN UT ES : S E CO N DS

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

7-]0



LOX

PUMP

INLET

TEMPERATURE,°

.

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F

T

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INLETTEMPERATURE,



The second burn start transient was satisfactory. The thrust buildup

was within the limits set by the engine manufacturer. Faster thrust

buildup to the 90 percent level, as compared to the acceptance test re-

sult, was observed on this flight. This buildup was similar to the

thrust buildup on previous flights. The PU valve was in the proper full

open (4.5 Engine Mixture Ratio [EMR]) position prior to the second start.

T he to ta l im pu lse from S T DV to S T D V + 2.5 seconds w as 7 9 7 ,335 N -s (1 7 9 ,248

Ib f-s). T h is was greater than the va lue o f 6 44,9 9 2 N -s (1 45 ,0 0 0 Ibf-s)obta ined during the sam e in terva l fo r the acceptance test.

S econd burn fuel lead appeared to fo llow the predicted pa ttern. E ven

thoug h the fuel in jector tem perature behaved in an erra tic m an ner, th e

fuel lead apparently resulted in satisfactory conditions as indica ted by

other m easurem ents and subsequent performance.

7 .7 S -IVB M AIN S T A G E P E R F O RM AN C EF O R S E C O ND B U R N

T he propu lsion reconstruction analysis showed that the stage performance

during m ainstage operation was satisfactory. A com parison of predicted

and actual performance of thrust, total flowrate, specific impulse, andm ixtu re ra tio ve rsus tim e is show n in F igu re 7 -7 . T able 7 -2 sho w s the

specific impulse, flowrates and mixture ratio deviations from the pre-dicted a t the 1 80 -second tim e slice. T h is tim e slice perfo rm ance is the

standardized altitude perform ance which is comparabl e to the first burnslice at 1 40 seconds.

T he 1 80 -second tim e slice perform an ce for second bu rn thrust was 0 .25

percent lower than predicted. S pecific im pulse performance for second

burn was 0 .30 percent h igher tha n predic ted.

S econd burn dura tion was 343.0 6 seconds from S T D V open, wh ich was 0 .6 5

second shorter than the predicted dura tion .

T he helium contro l system perform ed satisfactorily during second burn

m ainstage. T here w as little p ressure decay during the burn due to the

connection to the sta ge rep ressuriza tion system . H elium usage was esti-

mated from flowrates during eng ine opera tion . A pproxim ately 0 .358 kilo -gram (0.79 Ibm) was consumed during second burn.

Due to reports o f excessive v ibra tion during the fligh t, a specia l in-

vestigation has been undertaken concerning engine thrust variation in the1 8 to 1 9 hertz frequ en cy ran ge. S ince the P O G O effec t is a p oss ible

source of these vibrations, and it is known from prev ious experience that

the L O X pum p is responsive to P O G O d riv ing forces, investiga tion has been

concen tra ted on the L O X pum p. F requency M odu la tion (F M ) da ta su itable

fo r eva luation in the exp ected frequency ran ge, was eva luated for L O X

pump discharge pressure measurements for the AS-503 and AS-505 flights.O ther data fro m acceptance tests and other m easurem en ts were a lso eva lu -

ated. T h e data eva lu a ted so fa r h ave no t developed a pos itive in dica -

tion o f P O G O o r a positive corre la tion between thrust varia tions and other

flights or propellant conditions.

7-12



E

SPECIFI

II

P

LSE,-s/

g

T

RUST,

106

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A pp rop ria te data from past fligh ts and acceptance tests are being re-viewed in a deta iled manner in th is continu ing investiga tion .

7 .8 S -IVB S H U T D O WNT R A NS IE N T P E R F O RM A N CEF O R S E C O N D B UR N

T h e shutdown transient was satisfactory and agreed closely w ith the accept-ance test and predictions. T he to ta l cu to ff im pu lse to zero percen t o f

ra ted th rust was 21 0 ,6 50 N -s (47 ,356 Ibf-s). E C O w as in itia ted by a L V D C

velocity cu toff com m and. C uto ff occurred with the P U va lve in the nu ll

position.

7 .9 S -IV B S T A G E P R O PE L L A N T M AN A G E M EN T

O n A S -50 5 the P U system was opera ted in the open- loop m ode, w h ich m eansthe L O X flowrate is not con tro lled to insure s im ulta neous depletion of

propellants. T he P U system successfu lly accomplished _he requirements

associated with propellant loading.

A com parison o f prope llan t m ass va lues a t critica l fligh t events, as

determ ined by various analyses, is presented in T able 7 -3. T he best

estim ate fu ll loa d prop ellan t m asses were 0 .49 percent greater for L O X

and 0 .26 percen t less fo r L H2 than the predicted va lues. T h is devia tionwas well with in the requ ired loading accuracy.

T he th ird stage sta tistica l we ighted averag e m a sses a t ign ition were

1 6 5,57 3 and 1 32,6 0 0 k ilog ram s (36 5,0 25 and 29 2,332 Ibm ) fo r first and

second burn , respectively. T he cuto ff m asses were 1 33,830 and 6 2,450k ilog ram s (29 5,0 44 and 1 37 ,6 7 9 Ibm ) fo r first and second burn , respec-

tive ly. E xtra po la tio n of p ro pellan t level sensor data to dep letion ,

using th e propellan t flowrates to dep le tion , indica ted that a L O X dep le-

tion would have occurred approxim ately 1 0 .64 seconds after second burnvelocity cutoff.

T he P U va lve was positioned at nu ll fo r sta rt and rem ained there, as

programed, during first burn. The PU valve was commanded to the 4.5 EMRposition at 9 0 7 9 .3 seconds and rem ained there fo r 255.0 2 seconds. A t

9 334.3 seconds the va lve was com m anded to the nu ll position (approxi-

m ate ly 5.0 E M R ) and rem ained there th roughout the rem ainder o f theflight.

7 .1 0 S -IV B P R E S S U R IZ A T IO N S Y S T E M

7 .1 0 .1 S -lVB F uel P ressuriza tion S ystem

T he L H 2 pressuriza tion system operationa lly m et all eng ine perform ance

requ irem ents. T he L H2 pressuriza tion system indica ted acceptable per-

fo rm ance during prepressuriza tion, boost, first burn , coast phase, and

second burn. T he L H 2 tank pressurization com m and was received a t

-9 6 .41 seconds. T he pressurized signa l wa s received 1 3.1 seconds la ter.

7-14



Table 7-2. S-IVB Steady State Performance - Second Burn (ESC +180-Second

Time S lice at Standard Altitude Conditions)

PARAMETER PREDICTED SECONDBURN FLIGHT PERCENTEVIATION

RECONSTRUCTION DEVIATION FROM PREDICTED

ThrustN 9,129,304 9,106,040 -23,264 -0.25

(ibf) (205,235) (204,712) (-523)

S pecific ImpulseN-s/kg 4201 4214 13 0.30(Ibf-s/Ibm) (428.4) (429.7) (I.3)

L O X F lo wra te

kg/s 180.91 180.23 -0.68 -0.35(Ibm/s) (398.83) (397.34) (-I.49)

F uel F lowrate

kg/s 36.38 35.86 -0.53 -1.4(Ibm/s) (80.21) (79.05) (-1.16)

E ngine M ixtureRatio

(LOX/Fuel) 4.972 5.026 0.054 1.09

T able 7 -3. S -IVB S tage P ropellan t M ass H istory

PREDICTED PU INDICATED PU VOLUMETRIC FLOWINTEGRAL BEST ESTIMATEEVENT UNITS (CORRECTED)

LOX LH2 LOX LH2 LOX LM2 LOX LH2 LOX LH2

S-IC Liftoff kg 86,705 I9,73] 86,848 19,680 87,478 19,752 86,796 19,605 87,130 19,681

(Ibm)(19],]52) (43,500) (]9],466){43,386) (192,856 (43,545) (191,351)(43,222) :192,08g)(43,388)

First Ignition(ESC) k 986,705 19,731 86,844 19,671 87,443 19,750 86,796 lg,605 87,130 19,680(Ibm)(191,158) (43,500) (]91,458)(43,367) (192,778)(43,542) (191,35])(43,222) i19R,089) (43,388)

FirstCutoff (Eco) k 9 60,487 14,455 60,465 14,245 60,828 14,313 60,402 14,265 60,728 14,317(ibm) (133,350) (31,868) (133,302) (31,405) (134,108) (31,555) (133,164) (31,450) 133,883) (31,564)

Second Ignition (ESCl k B 60,360 13,177 60,274 13,142 60,681 13,210 60,256 13,162 60,54] 13,206(Ibm_ (133,072) (R9,051) (132,882) (28,973) (133,779) (29,123) (132,84l) (29,018) 133,471) (29,116)

Second Cutoff (ECO) kD 2248 918 2448 992 2420 g77 2410 995 2424 999(Ibm) (4957) (2025) (5396 (2]86) (5336) (2153) (5314) (2194) (5344) (2204)

Following the termination of prepressurization, the ullage pressure

reached relief conditions, approximately 21.8 N/cm2 (31.6 psia), and re-

mained just below this level at 21.7 N/cm2 (31.5 psia) until liftoff, as

shown in Figure 7-8. A small ullage pressure collapse occurred during

the first 20 seconds of boost and was followed by a return to the relief

level at 45 seconds due to self pressurization.

During first burn, the average pressurization flowrate was approximately

0.33 kg/s (0.72 Ibm/s) providing a total flow of 47.7 kilograms (I05 Ibm).

Ullage pressure was at the relief level throughout the burn, as predicted.

7-15



During the 02/H2 burner repressurization period, the LH2 tank was pres-surized from 13.3 to 20.8 N/cmz (19.3 to 30.2 psia). The LH2 ullagepressure was 21.7 N/cm2 (31.5 psia) at second burn ESC as shown in Figure

7 -9 . A pp ro xim ate ly 1 2.7 k ilog ra m s (28.0 Ibm ) o f he lium w ere used in th e

repressurization operation. T he average second burn pressurization flow-

rate was 0.30 and 0.32 kg/s (0.67 and 0.71 Ibm/s) for 4.5 and 5.0 EMR,respectively. At step pressurization the flowrate increased to 0.52 kg/s(1.14 Ibm/s). This provided a total" flow of 122 kilograms (268 Ibm)

during second burn. Significant venting during second burn occurred atsecond E S C + 280 seconds when step pressurization was in itia ted. T h is be-

havior was as predicted.

T he am bien t rep re_suriza tion system was used to repressu rize the tan k from

11.6 to 14.3 N/cmL (16.8 to 20.8 psia) for the LH2 lead experiment. Therepressurization was satisfactory.

S _7 P RE PR ES SUR IZ AT IO N IN IT IA TE D

S _7 F IR S T E S C, 553.6 0

_7 F IR S T E CO , 7 0 3.7 6R _7 R EP R ES SU RIZ AT IO N IN IT IA TE D

3o I ' I !, i • -40

E

u25 , I '_

_ , 30_

':=- 15 : _'-,, I I o_

MINIMUMPREDICTED ---:l'''=_ --...

.J

_j , ....I

m 10 _

e4 5 c,J

-J

0 , , 0• 0 1 2 5 6 7 8 9

R A N GE T IM E, I0 0 0 S E CO N DS

0:00:00 0:30:00 1:00:00 1:30:00 2:00:00 2:30:00

R A N GE T IM E , H O UR S :M IN UT ES :S E CO N DS

F igure 7 -8. S - IV B L H 2 U llage P ressure - F irst B urn and P ark ing O rbit

7-16



m

LH

U

AGE

N

cm

LH

U

AGE

N

cm2

_

_

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LHU

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Upa

,



T he L H2 pum p in le t N et P ositive S uction P ressure (N P S P ) was ca lcu la ted

from the pum p interface tem peratu re and to ta l p ressure. T hese va lues in-

dicated that the NPSP at first burn ESC was 10.5 N/cm2 (15.2 psid). Atthe minimum point, the NPSPwas 4.2 N/cm2 (6.1 psid) above the requiredpressure. T hrou ghout the bu rn , the N P S P satis factorily agreed with the

predicted value. The NPSP at second burn ESC was 5.2 N/cm2 (7.6 psid)

which was 1.8 N/cm2 (2.6 psid) above the required pressure. Figures 7-10and 7 -1 1 sum m arize the fue l pum p in let conditions for first and secondbu rn s, resp ectively.

7 .1 0 .2 S -IVB L O X P ressurization S ystem

L O X tank prepressurization was in itiated at -1 6 7 seconds and increased

the LOX tank ullage pressure from ambient to 28.3 N/cm2 (41.1 psia) within18 seconds as shown in Figure 7-12. Three makeup cycles were required tom ainta in th e L O X tan k u llage pressure before the u llage tem perature sta-

bilized. A t -9 7 seconds the L O X tank u llage pressure increased from 27 .4

to 28.5 N/cm2 (39.8 to 41.4 psia) due to fuel tank prepressurization, LOX

tank ven t purge and L O X pressure sense line purge. T he u llage pressureincreased steadily to 29.5 N/cm2 (42.9 psia) just before liftoff.

D uring S -lC boost there was a rela tively m oderate u lla ge pressure decay

caused by an accelera tion effect and tem perature decrease.

N o m akeup cycles occu rred un til an inh ibit was rem oved, app ro xim ate ly 50

secon ds befo re E S C . A t tha t tim e, o ne m akeup cyc le o ccurred. T h e L O Xtank ullage pressure was 27.5 N/cm2 (40.0 psia) at first ESC.

During first burn , th ree over-contro l cycles were in itia ted, as comparedto the predicted one cycle. T he L O X tank pressurization flowra te varia-

tion was 0.118 to 0.158 kg/s (0.26 to 0.35 Ibm/s) during under-control

system operation . T h is varia tion is norm al because th e bypass orificein le t tem perature changes as it fo llows the co ld helium sphere tem pera-

ture. Heat exchanger perform ance during first burn was satisfactory.

R epressurization of the L O X tank prio r to second burn was not required.

The tank ullage pressure was 27.5 N/cm2 (39.9 psia) at second ESC, satis-fying the requ irem en ts as sho wn in F igure 7 -1 3.

P ressurization system performance during second burn was satisfactory, and

had the sam e characteristics noted during first burn . A s predicted, there

were no over-control cycles. Flowrate varied between 0.25 and 0.31 kg/s(0.36 to 0.45 Ibm/s). Heat exchanger performance was satisfactory.

The LOX NPSP calculated at the interface was 19.1N/cm 2 (27.8 psid) atfirst burn E S C . T he N P S P decreased after sta rt and reached a m in im um va lue

of 17.3 N/cm 2 (25.1 psid) at 93 seconds after ESC. This was 6.8 N/cm2!9 .9 psid) above the requ ired N P S P at tha t tim e .

7-18



1 5

- 2 0

E(.3 "i_

z 10 IACTUAL

_- "10

5 / REQUIRED z

-# #

0 0

j 30 "50 __j'_ ,=Zl-- l--0

-45__

uJ u 25 _CTUAL--' _-'_-- _ "40 _, "T_ZZ ZC.

_ - -35_,.,:EZc_ ............. s-__ 20 -_

"30 _

=_ _PREDICTED _-__ -25 -J_-15 I

24 -- FIRSTIESCI I I

ST-/"-418 _:.,. S_7 FIRST BURNSTDVOPEN

23-_7FIRSTCO_.1 L.U

"-420

J ,--_ IL

--J _ _ _ _ _JZ Z

o_ 21 "-422_

223 _

_ 21Z

-J 20 "-424-J0 20 40 60 80 lO0 120 140 160

TIME FROM FIR ST ESC , SECONDS

550 570 590 610 630 650 670 690 710

R A N G E TIM E, SECON DS

Figure 7-I0. S-IVB Fuel Pump Inlet Conditions - First Burn

7-19



S_7SECONDBURNSTDVOPEN _ SECONDECO1 5

- 2 0

jACTUAL -z 10 / ' D_

CL 12L

o4 5 _I

REQUIRED"

0 I

j 30 -50< g:o -45ocJ ACTAL

ZQ

jo_ ....... "PREDICTED -25 _-_J(__

15 I I_z 24 a.0 0

"-418_- 23

(__ t.u

L_J LU

22_ "-420_-

"J zz _

o_ 21 "-422_-

-_ 200 .-424_0 1O0 150 200 250 300 350 400

T IM E F R OM S ECO ND E SC, S ECO N DS

, _ i J i i i ,_' ,2:33:00 2:34:00 2:35:00 2:36:00 2:37:00 2:38:00 2:39:00 2:40:00

R A NG E T IM E , H OUR S : M IN UT ES :S EC ON DS

F igure 7 -1 1 , S -IVB F uel P um p In let Conditions - S econd B urn

7 - 2 0



,oxANREP EsNIT,ATEO

CTUAL3 0

"- __= ---._ _ . _-40

25

z 20 -30

PREDCTED

m -20 _LIA _

_ CL

1(

-1 05

0 0-I000 0 I000 2000 3000 4000 5000 6000 7000 8000 9000

R A NGE T IM E, S ECO NDS

0:00:00 0:30:00 I:00:00 1:30:00 2:00:00 2:30:00

R A NG E TIME, HOUR S:MINUTE S: SE C ONDS

F igure 7 -1 2. S - IV B L O X T ank U llage P ressure - F irst B urn and P ark ing O rbit

T he L O X pum p sta tic in terface pressure during first burn fo llowed the

cyclic trends o f the L O X tank u llage pressure. T he N P S P ca lcu la ted a t

the engine interface was 16.0 N/cm2 _(23.3 psid) at second burn ESC. Atall tim es during second burn , N P S P w as above the requ ired leve l. F igures7 -1 4 and 7 -1 5 sum m arize the L O X pum p conditions fo r the first and the

second burn, respectively.

T he co ld helium supp ly w as adequate to m eet a ll fligh t requ irem ents. A t

firs t burn E S C the co ld helium sph eres con ta in ed 20 0 k ilog ra m s (442 Ibm )o f h e liu m . A t; the en d of the first bu rn , th e he liu m m ass ha d decreased

to 1 7 6 k ilog ram s (388 Ibm ). A t second burn E S C the spheres con ta ined1 6 3.3 k ilog ram s (36 0 Ibm ) o f he lium . A t the e nd o f second burn the he lium

m ass had decreased to 9 9 k ilogram s (21 8 Ibm ). F igu re 7 -1 6 show s he lium

supply pressure history.

7 - 2 1





_

u%

S3

O

'

S3O

0

-

O

m"

"

-

m

m

oL

O0

m

-__Q_,

'73_

O_

9_ 0

O0

×z

(

_

9

0_NO1



25 36

c , , l

ACTUAL 32 -o_> i// "F_0 _..

c,_ "_ _ "''''"-_ "28

z -24

x 15 xC _ C

' -20 "J

I0 .... _j RE_UIED................ "16

35 "50.-J

o ACTUAL _-

°...1 m

PREDICTED "'- "40 z

25 _

20 •30

9 7

_94 '-290_°,-..av Zt.LJ

o_m ........... 295_"_x_: 91 o__o_OLLI X_-_}-- OL_I

-Jl--.

--300

880 50 I00 150 200 250 300 350 400

T IM E FROM SECON D ESC, SECOND S

I I I I I I I |

2:33:00 2:34:00 2:35:00 2:36:00 2:37:00 2:38:00 2:39:00 2:40:00


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

7-24



_;ZF IR S T E S C

_T F IR S T E COS _Z SE CO N D E SC_zS ECO N D E CO

S _zS TA RT CO LD HE L IUM DUM P N O . 2

2500

. 3 0 0 0

2000

c_ JE

_-1500z -2000_

1000

I000

0 .00 2 4 6 8 I0 12 14 16 18 20

R A N GE T IM E_ I0 0 0 S E CO NDS

, , _Z , _ , , ,0:00:00 I:00:00 2:00:00 3:00:00 4:00:00 5:00:00 6:00:00

RANGETIME, IIOURS:MINUTES:SECONDS

F igure 7 -1 6 . S -IVB C o ld H elium S upp ly H istory

7-25



7 .1 2 S -IVB A UXIL IA R Y P R O P UL S IO N S Y S T E M

The operations of the APS pressurization system was satisfactory with theexception of a helium leak in Module No. I. The leak started approximately

6 .5 hours after lifto ff and extended through loss o f da ta at 39 ,240 seconds

(1 0 :54 :0 0 ). T h e lea k ra te a t lo ss o f d a ta w a s a p p ro x im a te ly 327 8 S C C M(20 0 S CIM ). F igure 7 -1 7 presents M odules N o. 1 and 2 helium bottle m assa t th e tim e o f th e lea k. T h e a ttitu de co n tro l requ irem en ts fo r M o du lesN o. 1 and 2 after 21 ,6 0 0 seconds (6 :0 0 :0 0 ) were approxim ately equal.

The range of regulator outlet pressure, ullage pressure, propellant mani-fo ld pressure, and propellant temperature is presented in T able 7 -4.

- 0.90

0.40_

-9.39

0.35MODULENO. 2

- 0.30 _ -

_._. _.-_MODULE NO. l

-0.60

0.25 _

-0.50

0.20 %

•0.40

10,800 14,400 18,000 21,600 25,200 28,800 32,400 35,000 39.600

R AN GE T IM E, S EC ON DS

3:00:00 4:00:00 5;00:00 6:00:00 7:00:00 8:00:00 9:00:00 |0:00:00 ll:O0:O0I I I I I I I I I


F igure 7 -1 7 . S -IVB A P S Helium B ottle M ass

7 - 2 6



T hese pressure values were satisfactory and with in instrumentation accu-

racy of the required values of 133 to 140 N/cm2 (193 to 203 psia) forregulator outlet and 130 to 138 N/cm2 (188 to 200 psia) for ullage andm anifo ld pressure. However, temperature extrem es of the regu la tor during

the la tter portio n o f th e m iss io n ca used the M o du le N o. 1 va lu es to in -

crease approximately 3.4 N/cm2 (5 psia) and Module No. 2 values to de-crease 2.1N/_n 2 (3 psia). The regulator temperatures were in the sameapproximate range as the helium bottle temperatures presented in Figure7 -1 8. S ince th is regu la tor was not tem peratu re co m pensa ted, these pres-

su re trends were exp ected w ith the tem perature extrem es seen.

A ll eng ines p erform ed satis factorily. A tim e histo ry of A P S prope llan ts

fo r M odu les N o. 1 a nd 2 is p resen ted in F igu re 7 -1 9 . T ab le 7 -5 p resen tsthe APS oxidizer and fuel consumption at significant events during thefligh t. T able 7 -6 sum m arizes the A P S sta tus a t loss o f da ta .

7 .1 3 S - IVB P R O P E L L A N T L E A D E XP E R IM EN T A N D O R B IT A L S A F IN G O P E R A T IO N

A propellan t lead experim ent was perform ed after spacecraft and L una rM o du le (L M ) separa tion . L O X a nd L H 2 flow th rou gh the eng in e sim u la ted

the contingency restart preparation sequence. T h is contingency sequence,which cou ld be used in case of recircu la tion ch illdown system failure,

p rov ided data fo r eva lua ting the adequacy o f the m ethod. B efo re and after

th is experim ent, the stage high pressure system s were sa fed. T he thrustdeveloped during the experiment and subsequent LOX dump was utilized to

ensure tha t the spen t S -IVB stage w ou ld be p laced in so la r o rbit.

T he m anner and sequence in wh ich the experim ent and sa fing were perfo rm ed

are presented in F igure 7 -20 .

T abl 7 -4. S -IV B A P S P ropellant C onditions

PARAMETER MODULEO.1 MODULEO,2

FUEL OXIDIZER FUEL OXIDIZER

Ullage Pressure

N/cm2 131 to 137 128 to 137 128 to 132 124 to 126

(psia) (190to198) (186to198) (]85to192) (180to183)

Propellant Manifold

Pressure

N/cm2 128 to ]34 135 to 138 124 to 131 125 to 131

(psia) (186to194) (196to200) (180to190) (182to190)

Propellant Temps

(In P ro pe ll an t Con tr ol

Module)

°K 290to304 290to307 305to316 304to315

(°F) (62to87) (62to93) (90toII0) (87to107)

Regulator OutletPressure

N/cm2 130to 139 130 to 139 128 to 134 128to 134

(psia) (]88to202) (188to202) (186to]94) (186to194)

7-27



360

, ] 8 0

350 i_

S NO.2 ']60

_MODULE

3 4 0

330 _ // 140320 /I _ 120 u_

j -N310 100N

_. 300

2 90i - 60 _

'_280 I N

= =_ I_' _ r-MODULE NO. 1

_: 270 -'-._ . /60

%_ 0

250 """""

_" -. -2024 0

• -4 0230,i

I0,800 14,400 18,000 21,600 25,200 28,800 32,400 36,000 39,600

RANGE TINE, SECONDS

i I i I i i I ! j

3:00:00 4:00:00 5:00:00 6:00:00 7:00:00 8:00:00 9:00:00 ]0:00:00 lI:O0:O0


Figure 7.-18. S-IVB APS Helium Bottle Temperature

7o13.1 LOX and LH 2 Lead Chilldown Experiment

The LOX and LH 2 chilldown experiment was successfully conducted as planned.

Preliminary evaluations indicate that propellant tank repressurizations

were within the limits predicted for the experiment and that the data re-

ceived, with appropriate analysis and interpretation, will provide chill-

down criteria for contingency restart procedures. The main LOX valve was

opened at 17,301 seconds and closed at 17,310 seconds, resulting in a LOX

lead time of 9 seconds. The main fuel valve was opened at 17,410 seconds

and closed at 17,459 seconds, resulting in a fuel lead time of 49 seconds.

The data received from this experiment have been evaluated from a "first-

look" standpoint and are presented in Figures 7-21 through 7-23.

LOX pump inlet conditions are presented in Figure 7-21. The data indicated

that the LOX pump inlet temperature went off-scale low 4 seconds after the

MOV opened and came back on-scale 49 seconds after the MOV opened; this

was 40 seconds after the MOV had closed. As shown in the figure, LOX pump

inlet temperature was satisfactory for engine start at the end of an

8-second fuel lead.

7-28



M O D UL E N O . 1I -220

S _T AP P RO X S TA R T O F H E L EA K

oOXIDIZER I

FUELI -200- P RE DIC TE D

BO Q C'-'qQ 180

16 0

14o

rnO] 0 \ ]zo_

(_ 1oo

<,40 u::L. ,-

5o _rn

2040

20

o o

MODULENO. 2 220

_, 200C ® E) '180

L_ .16o

®

"-3 _"_ _ ,140d

:_60 .]k,

_ \ .12o

'-- _ _ "lO0\_ _----_ -80 o

- 6 0

\20 ,40

-20

O .0

3600 7200 10,800 14,400 18,000 21,600 25,200RANGE TIME,SECONDS

I I I , , I I I I _ I

0:00:00 I:DO:O0 2:00:00 3:00:00 4:00:00 5:00:00 6:00:00 7:00:00


Figure 7-19. S-IVB APS Propellants Remaining Versus R ange Time,Module No. l and Module No. 2

7-29





Table 7-6. S-IVB Helium Bottle Conditions

MODULEO.1 MODULEO.2

CONDITIONST CONDITIONSTINITIAL LOSSOFDATA INITIAL LOSSOFDATA

PARAMETER CONDITIONS 39,000 SEC CONDITIONS 39,000 SEC

Pressure

N/cm2 2137 663 2137 1604

(psia) (3100) (961) (3]00) (2327)

Temperature

°K 305 237 304 351

(°F) (90) (-36) (87) (172)

Mass

k_ 0.4654 0.197 0.467 0.314

(ibm) (1.026) (0.435) (1.030) (0.692)

Usage

kg 0.268 0.153

(Ibm) (0.591) (0.338)

T he conditions a t the fue l pum p in le t a re presen ted in F igure 7 -23. F uel

m ea surem ents obta ined in dica te th at a ll- liqu id flow was present a t thep um p in le t 3 seco nds after fue l lea d sta rt. I t is a lso ind ica ted tha t

pump in let conditions remained substantia lly constant during the rem ainder

of the 49 -second fue l lead period. A s shown in the figu re , satisfactorysta rt conditions are projected for a norm a l resta rt L O X tank pressureconditio n and a n 8-second fuel lead.

T hrust chamber condition ing is depicted by the fuel in jection temperature

versus tim e curve in F igure 7 -23. T he m easurem ents indica te tha t the in -jection temperature chilldown characteristic demonstrated was within thepredic ted ran ge . H ow ever, it is n o t conc luded at th is tim e th a t a sa tis-

facto ry condition wou ld have existed if the fue l tank had been repressur-

ized to a norm a l resta rt p ressure leve l. T h is issue is clouded by erra tic

beh avior o f the fuel in jection tem perature m ea surem ent. T he m ost appro-priate adjustments have been made and are reflected in Figure 7-23; how-

ever, the response of th is m easurem ent is still under investiga tion .

7 .1 3.2 L O X T ank A m bient R epressuriza tion

A m bient helium repressuriza tion of the L O X tank in prepara tion for the

propellant lead experiment and L O X dump, was satisfactorily accomplished.

R epressuriza tion was in itia ted at approximately 1 7 ,1 53 seconds and was

term inated 20 2 seco nds la ter. He lium sup ply pressure dropp ed from 1 9 6 0

to 90 N/cm2 (2840 to 130 psia) and approximately 6.4 kilbgrams (14.2 Ibm)o f h e lium w ere a dded to the ta_k u llage . T he u llage pressure on ly in -creased from 17.5 to 20.4 N/cmL (25,4 to 29.7 psia) because of the largeullage volume.

7-31





7 .1 3.5 L O X T ank Dum p and S a fing

Imm edia te ly fo llowing second burn cuto ff, a program ed 1 50 -second vent

reduced LOX tank ullage pressure from 26.7 to 13.1N/cm2 (38.8 to 19.0

psia ) as shown in F igu re 7 -1 3. Da ta leve ls were as expected w ith 44.9

k ilog ra m s (9 9 Ibm ) o f h e liu m an d 7 2.5 k ilo g ram s (1 6 0 Ibm ) o f G O X be in gven ted overboard. A s indica ted in F igure 7 -1 3, the ullage pressure then

rose gradually due to self-pressurization, to 17.4 N/cm2 (25.3 psia) atthe initiation of ambient repressurization. R epressurization raised the

ullage pressure to 20.5 N/cm2 (29.7 psia).

T he L O X tank dum p was in itia ted a t 1 7 ,6 6 5.7 9 seconds and was satisfacto rily

accomplished. A steady-state liquid flow of 0.0260 m3/s (411 gpm) wasreached within 7 seconds.

A pproxim ately 55 seconds after dump in itia tion , the measured L O X flowrateshowed a sudden increase indica ting that gas ingestion had begun. S hortly

thereafter, the L O X ulla ge pressure bega n decreasing at a greater ra te .Ca lculations indicate the L O X residual, approximately 220 3 kilograms

(487 0 Ibm ), was dum ped with in 1 9 4 seconds T he tank p ressu re had decayed

to 13.3 N/cm2 (19.3 psia) at this time. Ullage gases continued to be

dumped until the programed term ination .

L O X dum p ended at 1 7 ,9 56 seconds as schedu led by closu re o f the M O V . A

stea dy-sta te L O X dum p th ru st o f 4340 N ew ton s (9 7 5 Ibf) w as o bta ined. T he

tota l im pu lse befo re M O V closure was 40 9 ,7 82 N -s (9 2,1 23 Ibf-s), resu ltingin a calculated velocity increase of 25.4 m/s (83.2 ft/s). Figure 7-24show s th e L O X flo w ra te during dum p an d th e m a ss of liqu id an d ga s in th e

oxidizer tank. F igure 7 -24 shows L O X u llage pressure and the L O X dum p

thrust p roduced. T he predicted curves prov ided fo r the L O X flowrate and

dum p thrust correspond to the quantity o f L O X dum ped and the actua l u llage

pressure.

T hree seconds fo llowing term ination of L O X dum p, the L O X N P V valve was

o pen ed a nd rem a in ed op en fo r the du ra tion o f the m ission . L O X ta nkullage pressure decayed from 8.5 N/cm2 (12.3 psia) at 17,956 seconds tozero pressure at approximately 25,0 0 0 seconds.

7 .1 3.6 C o ld H eliu m D um p

Cold helium was dumped through the 02/H 2 burner heating coils and into

the L H 2 tank, and overboard through the tank vents.

T he co ld he lium sp heres w ere sa fed by th ree co ld he liu m du m ps. D um p N o . 1

was in itia ted at 9 57 2 seconds and was program ed to con tinue fo r approxi-

m ate ly 87 8 seconds as show n in F igure 7 -1 6 . During th is period, the

pressure decayed normally from 358.5 to 34.5 N/cm2 (520 to 50 psia).A pproxim ate ly 6 0 .4 k ilog ram s (1 1 3 Ibm ) o f helium was dum ped overboard.

Dum p N o. 2 w as in itia ted a t 1 3,1 51 seconds and was program ed to con tinue

fo r approxim ate ly 89 9 seconds as shown in F igure 7 -1 6 . During th is period,

the pressure decayed normally from 68.9 to 6.9 N/cm2 (I00 to I0 psia).

7-33



S TA RT O F L O X L EA D + 4 S EC O N DSS TA RT O F L O X L EA D + 49 S EC O N DSS TA R T O F F UE L L EA DE N D O F 49 S E CO N D SF UE L L E A DP RO JE CT EDF O R F U EL L EA D S TA R T + 8 S E CO N DSW IT H N O R M A L LO X T AN K R E ST AR T C O ND IT IO N S

L OX IN TE RF ACE S TA TIC P RE SS UR E, psia

28 32 36 40 44 48

lO0 l _ 'I I' i

• AS-503 FIRST STARTSATURATION / • AS-503 SECONDSTART

=_ START

/ '_ Iv/-- LIMITS-- -27698 /

/

Pd

° 96 / I -/ -282

94 I0 /-28S

I 0 Iz 92

× I " I ×0

I I -29__0 I I

QOFFSCALE -300

88 I,r18 20 22 24 26 28 30 32 34

LOX INTERFACESTATIC PRESSURE,N/cm2

F igure 7-21 . L O X P ump Inlet Chilldown E ffectiveness

A pproxim ately 1 2.7 kilogram s (28 Ibm ) of helium was dum ped overboard. A ninsignificant amount of helium was dumped overboard during the third coldhelium dump which was in itiated at 1 6 ,937 seconds and lasted 1 51 1 seconds.

7 .1 3.7 A m bient Helium Dum p

T he am bient helium remain ing in the L O X and fuel repress spheres was

dumped through the engine contro l helium regulator via the engine contro lsphere. T he fuel repress spheres pressure decay began at 1 7 ,9 65 secondsand lasted for 230 1 seconds. T he pressure decayed from 6 20 .1 to 7 5.8N/cm2 (900 to llO psia). The LOX repress spheres p_essure decay began at1 8,30 0 seconds and lasted 96 6 seconds. T he pressure decayed from 20 3.2

to 75.8 N/cm2 (295 to llO psia). The LOX and fuel repress spheres weresecured by term inating the engine control bottle dump.

7-34



LOX PUMPDISCHARGEPRESSURE, psia

20 24 28 32 36 4010 4

I ' ' ' ' ' -278

_L O X L E A D S T A R T + 0 S E C O N D SI02 i k._)END OF 9 SECONDLOX LEAD

(_)FUEI.EADSTART+8 SECONDS SATURATION__

('_PROJECTEDORFUELLEADSTART+8SECONDS I/ " -2_2 _vWITH NORMAL LOX TANK RESTARTCONDITIONS

1 1O0

-290

96 _

o / (_ 503 =0SECOND

BURN -29494

/ START

/ F LIMITS 50392 0FIRST -298

B U_

90 (_12 14 16 ]8 20 22 24 26 28 30

LOX PUMP DISCHARGE PRESSURE, N/cm2

Figure 7-22. LOX Pump Discharge Chilldown Effectiveness

7.13.8 Stage Pneumatic Control Sphere Safing

The stage pneumatic control sphere was safed by initiating the J-2 engine

pump purge and flowing helium through the pump seal cavities to atmos-

here. The stage pneumatic control sphere dump was initiated at 16,936

seconds and had a programed duration of 3600 seconds. The pressure de-

cayed normally from 2034 to 868 N/cm2 (2950 to 1260 psia). The safing

period satisfactorily reduced the potential energy in the sphere.

7.13.9 Engine Start Sphere Safing

The engine start sphere was safed during a 150-second period at approxi-

mately 9553 seconds. Safing was accomplished by opening the sphere vent

valve. Pressure was decreased from 785.5 to 13.8 N/cm2 Ill40 to 20 psia)

with 1.5 kilograms (3.3 Ibm) of hydrogen being vented.

7.13.10 Engine Control Sphere Safing

The engine control sphere was safed beginning at 17,965 seconds and ending

at 19,266 seconds. The helium control solenoid was energized to vent

helium through the engine purge system. The pressure decayed from 620.I

to 75.8 N/cm2 (900 to llO psia). The ambient helium remaining in the

LOX and fuel repress spheres was also dumped via the engine control

sphere.

7-35





'_7ULLAGE GAS INGESTION STARTS

IOO0

4000

G O 0 _z

3000

_= 6oo=

2000

400

I000 200

0 0

' 30

20._

\_ ,25

] 5 I

,20

I0" _ .15

\17,656 17,756 17,856 17,956


t _ I I I

4:5_:10 4:55:50 4:57:30 4:59:10

R ANGE TIME, HOUR S:MI NUTES:SECONDS

F igure 7 -24. S -IV B L O X D um p (S hee t 1 o f 2)

7 - 3 7



LOX

MASS,

kg

LOX

FLOW_TE

,

kg/

4

/

=

N

_

_

x_

g

_

N

°

'

l

_

N

N

N

N

_

LOX

MASS,

Ib

LOX

FLOWRATE,

lbm/



S E CT IO N 8

HY DR AUL IC S YS TE MS

8.1 S U M M A R Y

T he stage hydrau lic system s perform ed satisfactorily on the S -IC , S -II,

and firs t burn an d co ast p hase o f the S -IV B stag e. D u ring th is pe riod

a ll param eters were w ith in specifica tio n lim its and there were no dev ia-

tions or anom alies. S ubsequent to th is tim e, during second burn and

translunar coast, there was a m inor p roblem with the engine driven hy-

drau lic pum p and an apparently unrela ted pro blem with the a uxilia ry hy-draulic pum p. S hortly after second burn sta rt com m and the eng ine driven

pum p output pressure sligh tly exceeded the compensator setting , but sys-

tem perform ance con tinued to be nom ina l during the burn . S om etim e during

the second burn the auxilia ry hydrau lic pum p p erform ance was degraded as

evidenced by system response after E ng ine C uto ff (E C O ) and during coast

phase activities. However, there was no indication of mission or pro-gram im pact due to th is anom aly.

8.2 S - IC H Y D R A U L IC S Y S T E M

A nalysis indicates that a ll servoactuators performed as com manded duridg

the fligh t, w ith a m axim um deflection equiva lent to 2.1 5 deg rees eng ine

g im ba l angle a t approxim ate ly 82 seconds. A ll o f the hydrau lic supp lypressures and tem peratu res were with in opera ting lim its with the excep-

tion o f eng ine N o. l c losing pressure. T h is m easurem ent sta rted to in -

crease unexpectedly a t 80 seconds as show n in F igure 8- I, and reached a

maximum of approximately 172 N/cm2 (250 psia) near the end of S-IC flight.T h is app arent increa se wa s due to instrum ent error.

8.3 S - II H Y D R A U L IC S Y S T E M

The S-II hydraulic system performance was normal throughout the flight.S ystem supply and return pressu res, reservo ir vo lum es, and system flu id

temperatures were within predicted ranges. Reservoir fluid temperatures

were close to the predicted ra te o f increase. A ll servoactua to rs re-sponded to com m ands with good precision , and fo rces acting on the actua -

to rs were well below the predic ted m axim u m .

8.4 S -IV B H Y D R A U L IC S Y S T E M (F IR S T B U R N )

The S-IVB hydraulic system performance was nominal throughout S-IC/S-IIboost a nd S -IVB first burn .

8 - I



_7 CE'CO .2000_70ECO

1 2 0 0 I

•1500

(xI ,_

300. _-

•1000:

ENGINEO.1-_ 500

,/_i_.z ...... L_iI_ 1111_ 1111 rlli vJ/fJ/I/] i//J r f/J, r/]/f "fff_iJJAfJffvJfJ_fJf "_JfF _ _ 7

00 20 40 60 80 I00 120 140 160

RANGETIME, SECONDS

F igure 8-I. S -IC E ngin e, V aT ve C losing P ressure

The supply pressure was nearly constant at 2500 N/cm2 (3630 psia), whichis within the allowable 2413 to 2517 N/cm2 (3500 to 3650 psia).

T he system internal flu id leakage was shared by the m ain engine driven

and auxilia ry pum ps during engine burn as characterized by a sligh t rise

in system pressure after ignition and the auxiliary pump motor current

drain of 32 am peres. T he auxiliary pum p, therefo re , was supp lying approxi-mately 25.2 cmJ/s (0.4 gpm) of the total leakage flowrate. The engine sup-plied 3.85 horsepower to the m ain pum p during the burn.

E ngine deflections were nom ina l th rou ghout first burn . T he actua tor po si-

tions were offset from nu ll during powered fligh t due to the disp lacem ent

o f the C ente r o f G rav ity (C G ) o ff th e veh ic le cen ter line , en g ine ins ta lla -

tion tolerances, thrust misalignment, and uncompensated gimbal clearancesand thrust structure compression effects.

8.5 S - IV B H Y DR A U L IC S Y ST E M (P A R KIN G O RB IT C O AS T P H A S E )

During the orbita l coast phase, two hydrau lic system therm a l cycles o f 48

seconds dura tion w ere program ed at 330 4 and 6 1 0 4 seconds. T he purpose of

these cycles is to distribu te h eat throughout the system by circu la ting

hydraulic fluid periodically.

A fter E C O th e pum p in le t o il tem p era tu re increa sed fro m 323°K (II9 °F ) to

a m axim um of 346 °K (1 6 4°F ) prio r to the firs t therm a l cyc le, w h ich w as

well w ith in the upper lim it o f 40 8°K (27 5°F ).

8-2



8 .6 S -IV B H Y D R A U L IC S Y S T E M (S E C O N D B U R N )

T he au xilia ry pum p was turned on during second burn prestart prepara tio ns

a t approxim ately 8848 seconds. S ystem op eration was norm al throug hout

this period. Shortly after engine start, system pressure increased from2502 to 2695 N/cm2 (3635 to 3770 psia). This pressure step exceeded the

pump compensator upper limit of 2508 N/cmL (3650 psia) as shown in Figure8-2. However, pump inlet and reservoir oil temperatures increased at thenominal rates of 5.2 and 2.0°K/min (9.4 and 3.6°F/min), respectively.E ngine deflection s were n om ina l th rou ghout the burn as shown in F igure8-3. T herefo re , th is 3 percen t excess in system pressure is no t consid-ered to be a problem .

S ystem leakage during second burn was furn ished by the engine driven

pum p. T he eng ine supp lied 4.85 horsepow er to drive the pum p during th isperiod.

8.7 S -IVB H Y DR A U L IC S Y S T E M (T R A N S L U N A R IN JE C T IO N C O AS T A N D P R O PE L L A N TD U M P )

Degraded performance of the auxiliary hydraulic pump was observed duringthe period beg inn ing with second burn E C O . Data indica ted tha t the ano-

maly orig ina ted during second burn. S ystem pressure decreased im media te lyafter E C O as shown in F igure 8-2, whereas norm a l operation pressure would

be m ainta ined until the auxilia ry pum p "O F F " com m and was g iven 3.8 seconds

later. Failure of the auxiliary pump motor amperage to rise during thisperiod after ECO further substantiates degraded pump performance as shownin F igure 8-4.

A third therm a l cycle , a t 1 2,7 49 seconds, tu rned the auxiliary pum p on

fo r 48 secon ds;. N o increase in system p ressure, accu m ula tor G N2 pressureor reservoir o il pressure was observed. A ft ba ttery N o. 2 m easurem ent

indica ted 1 7 am peres th roughout the cycle as com pared to a predicted valueof 38 to 42 am peres as shown in F igu re 8-5 . T he actua tor position m ea-surem ents shown in F igure 8-6 indicated tha t the actuators cen tered. It

requ ired 31 and 1 7 seconds, respective ly, to cen ter the pitch and yawactuato rs. T he eng ine driven pum p in let tem peratu re dropped to a m inim um

va lue o f 352°K (1 7 3°F ). T h is actua to r m otion and o il tem pera tu re decreaseconfirms some auxiliary pump output pressure.

W hen the auxilia ry hydrau lic pum p was activa ted for th e prop ellan t leadexperim ent and p assiva tion , system perform ance was very s im ila r to th at

of the third thermal cycle. However, a slight increase in reservoir oilpressure of 50 to 55 N/cm2 (72 to 80 psia), as shown in Figure 8-7, wasobserved. A lthoug h the system was not perform in g properly , eno ugh systempressure was m ain ta ined to center the actuators for p assiva tion as shownin F igure 8-8.

S ubsequent labora tory testing was accomplished by simula ting fa ilures that

cou ld have caused th is anom aly. O f the sim u la ted fa ilures, the pressure

8-3



com pen sator sprin g gu ide fa ilu re test p ro duced data closest to thatobserved during th e fligh t. T he pressure com pensator spring gu ide

has been replaced on A S -50 6 and A S -50 7 veh ic les. T here is no indica -

tion o f any m ission or p rogram im pact due to th is anom aly.

3 0 0 0

• 4 0 0 0

C_JE Go _

z Y

w" 2000 ACCUMULATORN PRESS 3000_u_

dJ_ O

1500 f 2000

1 0 0 0

2 O O J

A UXIL IA R Y H YDR A UL IC P U M P O N

_ S E C O N DE S CE CO N D E CO

150 PREDICTEDIMITS

_o _T__H:__ _ooo.... .......ESERVOIROIL PRESS

w IOO _"

w -I00cI .

50 - - '- - .:;-

0 .0

8600 8800 9000 9200 9400 600 9800


2:25:00 2:30:00 2:35:00 2:40:00

R ANGE TIME, HOURS:MINUTES:SECONDS

Figure 8-2. S-IVB Hydraulic System Pressure - Second Burn

8-4



2 . 5 0

2.00 _W_.IlJ

1.50Z

o 1.00

j0.50 __ Lt ,& ......

0.00o -_'_'_'_-__-"_- ............. , ""F-

"<-0.50 ...............F--

___AUXILIARY HYDRAULICPUMPON'<-I.OO

_ SECONDSC_- -1.50 t I

_ SECONDCO-2.00 l _ I

P R E DIC TE D L IM IT S-2.50

2 . 5 0

2 . 0 0

l .50

z 1.00 ---

0.500

0O0 "=rv

-o. 50 - l"" _" 1"" .... • ,4;_I- -

- I .oo

< -I .50

>- .j f

-2.00 _,0_ _ _--

- 2 . 5 0

8600 8700 8800 8900 9000 9100 9200 9300 9400 9500 9600

R AN GE TIM E, S ECO NDS

2:25:00 2:30:00 2:35:00 2:40:00

RANGE TIME, HOURS :MINUTES:SECONDS

F igure 8-3. S -IVB H ydrau lic S ystem A ctua tor P ositions - S econd B urn

8-5





40 0

'24038O

_-- --_-----_ /'---'-YDRAULIC PUMP INLET OIL ,200° TEMPERATURE °

u_360 "- ' u_

_340 -160

RESERVOIRILTEMPERATURE

80

_7 AUXILIARY HYDRAULICPUMPON

_7 AUXILIARY HYDRAULICPUMPOFFPR EDIC TE D L IMIT S

lO O

8O

2O

0

12,700 ]2,740 |2,780 12,820 12,860 12,900

RANGE TIME, SECONDS

I

3:32:00 3:33:00 3:34:00 3:35:00

R AN GE TIME, HOUR S:MINUT ES :S EC ON DS

F igure 8-5. S -lVB A uxilia ry H ydraulic P um p P erfo rm ance - C oast P hase

and T hird T herm al Cycle

8-7



z0

IIIII

--

-2 _-ZAUXILIARYYDRAULICUMPON__

2 _ AUXILIARYYDRAULICUMPOFF--

PREDICTED LIMITS

o IIIIlO

F-

_ 0

__/--

- 2

12.700 I2,740 12,780 12,820 )2,860 12,900

R AN GE TIME , S EC ON DS

, _7 , _7 . j3:32:00 3:33:00 3:34:00 3:35:00


Figure 8-6. S-IVB Hydraulic System A ctuator Positions - C oast Phaseand T hird T hermal C ycle

8 - 8



1 5 0

-200

1 2 5

' 1 6 0

I00oJ "_E RE_;ERVOIRIL PRESSUREu '120

z 75

:_ bLI

u_ 50 _'_'_'_ w,,_,,_ ,_'>_ _ _ _ _ _ -80 ,,'_"

25 '40

O 0

_7 AUXILIARYHYDRAULICPUMP ON

x_7AUXILIARYHYDRAULICPUMPOFF

PREDICTED LIMITS

60.

----0 , . ___

I ,17,000 17,600 18,200 18,800 19,400 20,000

RANGETIME,SECONDS

_7, , , _ , _4:50:00 5:00:00 5:IO:O0 5:20:00 5:30:00

R AN GE T IME, HOUR S:MINUT ES :S EC ON DS

Figure 8-7. S-IVB Auxiliary Hydraulic Pump Performance - Coast Phaseand Passivation

8 - 9



0

0 i

I- -

-2 X_7 AUXILIARYHYDRAULICPUMPON

2_- _7 AUXILIARYHYDRAULICPUMPOFF-

PR ED IC TE D L IM IT S

"_ 1

0

o L,, __,_,.J_.eM_ _ _ ;_. _ ..__- 0

I" -

- 2

17,000 17,600 18,200 18,800 19,400 20,000

_NGE TIME, SECONDS

X_, , , _, _7_4:50:00 5:00:00 5:10:00 5:20:00 5:30:00


Figure 8-8. S-IVB Hydraulic System Actuator Positions - Coast Phase

a nd P as si va ti on

8-10



S E CT IO N 9

STR UCTUR ES

9 .1 S U M M A R Y

T he structura l loads and dynam ic env ironm ents exp erienced by the A S -50 5

launch veh ic le were well w ith in the veh icle structura l capability. T h e

veh icle loads resu lting from rigid-body and dynam ic longitudinal load and

bending m om ent were well be low lim it design va lues.

T he m axim um bending m om ent condition , 9 .9 x 1 0 6 N -m (88 x 1 0 6 Ibf- in .),was experienced at 84.6 seconds. T he m axim um longitudina l loads on theS -IC thrust structure, fuel tank, and in tertank were exp erienced at 1 35.2

secon ds, C en ter E n g ine C u to ff (C E C O ). O n a ll the veh ic le stru ctu re above

the in tertank , the m axim um lon g itudin a l loads were exp erienced at 1 6 1 .6seconds, O utboard E ng ine C uto ff (O E C O ), a t the m axim _ long itudina l

accelera tion of 3.9 g .

Vehicle dynamic characteristics genera lly followed the preflight predic-tions. There was no evidence of coupled structure/propulsion system in-

stab ility (P O G O ) durin g S - IC , S - II , o r S -IV B pow ered fliqh ts. T he early

S -II stage center eng ine shutdown successfu lly elim inated the low-fre-

qu ency (1 6 to 1 9 hertz) oscilla tion s tha t w ere experien ced on A S -50 3 a ndAS-504.

D uring S - IV B firs t and seco nd burns, m ild low -frequency (1 2 to 1 9 hertz)oscilla tions were exp erienced with the m a xim um am plitu de o f ± 0 .30 g re-

corded by the g im bal block long itu dina l acceierom eter. Durin g the last7 0 seconds of second burn , the A po llo I0 astronau ts reported (in real

tim e) tha t h igher frequen cy oscilla tions were superim posed on ithe low-frequency oscillations. T hese vibra tions are, however, well w ithin

the structura l design capability.

T he A S -50 5 vehicle structure, component, and engine vibration measurements

were, in g enera l, w ith in the envelo pes established by prev ious fligh t data .

9 .2 T O TA L VE H IC L E S T R U C T U R E S E VA L U A T IO N

9 .2.1 L ong itudinal L oads

T he A S -50 5 veh icle liftoff occurred nom ina lly a t a steady-sta te accelera-

tion of approxim ate ly 1 .2 g . T ransien ts due to th rust bu ildup and re lease

9 - I



resulted in peak longitudinal dynamic accelerations, measured on the

outboard and center engine thrust pads, of ±0.5 g and ±I.05 g,

respectively. These responses were less than 20 percent of the 3-sigma

95-percent confidence design level.

The AS-505 slow-release rod force displacement characteristics are com-

pared to the previous flight data in. Figure 9-I. The higher release rod

forces on AS-504 and AS-505 are believed due to less greasing of the rods.

The longitudinal loads that existed at the time of maximum aerodynamic

loading (84.6 seconds) are shown in Figure 9-2. There were no discern-

ible longitudinal dynamics at this time. The steady-state longitudinal

acceleration of 2.19 g and the corresponding axial loads experiencedwere as expected.

The maximum longitudinal loads on the S-IC thrust structure, fuel tank,

and intertank occurred at 135.2 seconds (CECO) at a longitudinal accelera-

tion of 3.67 g. (See Figure 9-2). The maximum longitudinal loads on all

vehicle structure above the S-IC intertank occurred at 161.6 seconds (OECO)at an acceleration of 3.9 g. The thrust cutoff transients experienced on

the AS-505 vehicle are shown in Figure 9-3 and are essentially identicalwith those of the AS-504 vehicle.

DIS PL AC EME NT , in.

0 1 2 3 4 5 6 7 8 9

500 l ! I

.................S-501 RELEASE ROD FORCE FROM MEASURED DATA

AS-502 R ELEASE R OD FOR CE FR OM MEASURED DATA

....... AS-5O 3 RELE ASE R OD F OR CE FR OM MEASUR ED DATA

AS-SO4 RELEASE ROD FORCE FROM MEASURED DATA .lOO

400 1 AS-505RELEASEOD FORCEFROMMEASUREDDATA--

__ Y///z//////////_/ALLOWABLEISPERSION BAND--"-\ .ao

3oo \

_ d200

I00

=• o0 O.O5 O.lO 0.I5 0.20

DIS PL AC EME NT , m

Figure 9-I. R elease R od Force - Displacement Curves

9-2



VEHICLESTATION, in.

I I / I I3000 2000 1000 0

VE HICL E S TA IIO N, m

90 80 70 60 50 40 30 20 ]0 0

I I lt : 84.6 S EC

ACCELER ATION = 2.1g G50

t = 135.2 SEC

.... ACCELER ATION = 3.67 G-IO

t = 161.6 SEC

----- ACCE LER ATION : 3.9 G40

30 , __ /L--F /

3 i -6

_.-s._=-z_-__-. <_2O

-4 _

lO -2

, _,_

Figure 9-21. Longitudinal Loads at Maximum Bending Moment, Center

Engine Cutoff, and Outboard Engine Cutoff

9.2.2 Bending Moments

The lateral loads experienced during thrust buildup and release were much

lower than design because of the favorable winds experienced during launch.

The wind speed at launchwas low, 8.2 m/s (16 knots),at the 18.3-meter(60-ft) level. The comparable launch vehicle and spacecraft peak redline

wind is 18.9 m/s (36.8knots) and 14.4 m/s (28 knots),respectively.

The inflight winds that existed during the maximum aerodynamic loading

phaseof the flightwere measuredat 42 m/s (81.6knots) at 14 kilometers(45,932 ft) altitude. T hese winds were approximately one-half the velo-

city of those encountered during the AS-504 flight. However, the trajec-

tory for AS-505 was not wind biased (for the first time for Saturn V

flights) and, as a result, the maximum bending moments experienced by

AS-505 were about the same as for AS-504, about 40 percent of design value.

As shown in Fiqure 9-4, the maximum bending moment of 9.9 x lO6 N-m (88

x lO6 Ibf-in.) was experienced on the S-IC LOX tank at 84.6 seconds. Load

9 - 3



CO,MANUOUUEE VEHCLESTATION1 2

VEHICLETATION7.2,. (INCHES) (METERS)_"st° (3B28..) IN

4 AS-504

2 -_ _ r-o.B -- 4245.4 107.8

O I "P_ Jv Iv 1" IV

AS-SOS2 I -o3890.3 98.8

I N ST R UM E N T U N IT

-- _ 3781.0 96.0VEHICLE STATION82,7 m

(3250i,,)

4.0 I _ -- _ 3594.6 91.3

--_ -- 3340.3 84.8

3258.6 82.83.0 _ 3222.6 81.9

3100.678.8.0 _ 2962.9 75.22832.0 71.9

_ _A_l_/ I ]t-AS-5O5 MEASURED DATA] 2645.2702.40 67.68'11.0-- .- rAS-SOMEiSUREOATCI/"_l 2519.0 64.0I i (-.... " -- 2387.0 60.6

- T A o R o o u , -- i i ," [ "

-].

_62 IB3 I64 1848.0 46.9

RANGEIME,ECONOS 1716.0 43.6

OUTBOARD ENGINE GIMBALBLOCK ]541.0 39,l

5.0 VEHICLETATION.5m ]401.0 35.6

(uTon.)-----_AT#.0 DROP

, 912.0 23.1

-- 612.0 15.520 I

LIJ

_.o 1 -- 225.0 5.7

-1.

160 161 162 163 164 165

RANGE TIME, SECONDS

F igure 9 -3 . L on gitudin al S tructural D ynamic R espon se D ue toOutboard Engine Cutoff

9 _ 4



VEHICLESTATION,in.3000 2000 lO00 0

I I I !

VEHICLESTATION, m90 80 70 60 50 40 30 20 ]0 0

--r r T--T---------I" I I" T T--

3 O 0t : 84.6 SECcl = 4.06DEG

_;AVG. = 0.71 DEG

CALCULATEDQUASI-STEADYLOADS _-----',_DESIGN LOADS

(,)TRAINGAUGE LOADS i_- _" _"

"T 200 / NORMALLOAD \

II F FACTOR \\ O.08

// \\ g0.06_

lOO /" ', 0.04._

j.i "/_ • 0.02

F igure 9 -4. M axim um B ending M om ent N ear M ax Q

computations are based upon measured inflight parameters such as thrust,

gimbal angle, angle-of-attack, dynamic pressure, and accelerations. The

bending moment values indicated by circles were derived from measured

strain gage data.

9.2.3 Vehicle Dynamic Characteristics

9.2.3.1 Longitudinal Dynamic Characteristics. The predicted first

longitudinal mode frequencies were present throughout the AS-505 S-IC

boost phase. (See Figure 9-5.) The measured frequencies agree well

with the analytical predictions. The frequencies are determined byspectral analysis using 5-second time slices.

The S-IC CECO transients were comparable in amplitude and frequency to

those observed on AS-504. The amplitudes were slightly lower initially

on AS-505, but decayed as slowly as on AS-504, indicating that vehicle

damping in this mode was again low. The data of Figure 9-6 show peak

amplitude of first mode oscillations versus body station for 135 through

9-5



LEGEND

I0

COMMANDODULE _ALYSlS FIRST MODEMEASUREDIRSTMODE 0

i , _ALYSISTANKBULGINGMODE•

I r ! MEASUREDANKBULGINGMODEJ

i i

N

f

4q_ _ _---- _-_-

i i

i I2 I i _ J i I1 i !

I _ ! ir I PI ; f I

, ; i I

0 1 i i I_ i _ _

0 20 40 60 80 I00 120 140 160

_N GE T IM E, S E CO NDS

F igure 9 -5. F irst Longitudinal M odal F requencies During S -ICP owered F lig ht



4000 100F IR S T LO NGIT UDIN AL M ODE

1 38 S E C 5 .2 H E R T Z

I3600 90

3200 80

2800 70 D

E

z oo 2400

_ 60I--- ,_

I-- c_

UJ

2000 _ 50

1600 40

120(} 30

800 20

400 I0

0 0

-O.5 0 0.5

A CCE LE RA TIO N , G peak

Figure 9-6 Peak Amplitudesof VehicleFirstLongitudinalModefor AS -504 and AS-505

9 - 7



1 38 seconds tim e slice. T he am plitudes o f m easurem ents on both A S -50 4

and A S -50 5 fligh ts are sho w n as w e ll a s a fit o f th e p redicted firs t

veh icle longitudina l mode through the data po in ts.

T he S -IC O E C O transien ts that were experienced by the A S -50 5 veh iclewere nominal and were nearly identical to those experienced on AS-504.

The S-IC/S-II separation dynamics were as expected. A maximum -0.6Gpeak acceleration was measured in the command module as compared to a

-0.8 Gpeak acceleration on AS-504 (See Figure 9-3).

The early S-If stage center engine shutdown successfully eliminated the

low-frequency (16 to 19 hertz) oscillations that were experienced on the

AS-503 and AS-504 flights. As shown in Figure 9-7, the AS-505 center

engine crossbeam response levels after S-II CECO were generally below

the readable threshold level of ±0.3 g as compared to the ±12 g ampli-

tudes on AS-504. The maximum amplitude measured on AS-505 was ±2.0 g

at S-II CECO, and the maximum sustained response was about ±l.O g at

approximately 294 seconds.

The most significant structural responses during the AS-505 flight

occurred during S-IVB first and second burns. Low-frequency (12 to 19

hertz) oscillations were experienced during both burns. During first

burn, a 19-hertz sinusoidal oscillation began on the J-2 engine gimbal

block (AOl2) at about 592 seconds. The oscillation reached a maximum

of ±0.30 g at 620 seconds, and decayed to negligible vibration by 639

seconds. Both the oxidizer pump discharge pressure (DOO9) and the main

LH 2 injector pressure (DO04) showed increases in 19-hertz oscillations

during this time period. Maximum pressure variations at 19 hertz were

±3.03 N/cm2 (±4.4 psia) for DO09 and ±0.9 N/cm2 (±1.3 psia) for DOO4.

During S-IVB second burn, the Apollo lO astronauts reported (in real time

at 9486 seconds) experiencing high-frequency vibrations. Recapping later(at I0,415 seconds), they reported lateral and longitudinal low-frequency

oscillations throughout first and second burns, and compared the flight

to a rough-running Titan; they reconfirmed a definite shift to a high

frequency superimposed upon the low frequency during second burn. The

high frequency was estimated to be approximately 20 hertz. Another com-

ment made at this time was " ... we were sweating it all the way, but

it shut down right on time." The comments from crew debriefing meetings

since mission completion have not reflected the same severity as in real

time and in inflight recaps; however, they confirmed that the S-IVB

second burn high-frequency oscillations were audible and could be felt

in the structure of the command module.

The flight measurements show a correlation with the astronauts reports.

Several measurements detected the sudden shift to high-frequency (45-

hertz) oscillations at 9481.8 seconds. These oscillations continued un-

til S-IVB engine cutoff (second ECO). The amplitudes for many measure-

ments, although low, also show a definite increase at this time. The maximum

9-8



I

E5 LONGITUDINALVIBRATIONBEAM (E363-206)

AS-504

10-

_ O- _

d -IO

-20

ZO

AS-505

_ lo>

-I0

-2 O

LONGITUDINALVIBRATIONLOXSUMP(E362-206)

20 '

AS-50410 -

-io-

N -2o

_ 2oAS-505I0

-lO

-20

505 506 507 508 509

RANGETIME, SECONDS

Figure 9-7. Comparisonof AS-504and AS-505 S-IIStage LowF r equ en cy O scillatio n s

vibration levels at 45 hertz were m easured by the S -IVB forward skirt

p itch and yaw accelerom eters as shown in F igure 9 -8. T he pitch measure-ment E 0 99 indicated a maximum of +_0 .58 g.

T he L H2 step pressurization event occurred at 9479 .2 seconds, which was2.6 seconds before the vibration level increase. F ollowing the step

pressurization,ltheon PropulsiveVent (NPV)nozzlepressuresincreased

as expected. At 9481.3 seconds, the NPV pressures began oscillations

at about_+1.4N/cm2 (+2 psia). It couldnot be determinedif the pres-sure was oscillating at 45 hertz because of the low sample rate in the

pressure measurements (D183 and D184); however, it does appear that the

9 - 9





T able 9 -- I. S -IVB S tage L ow-F requency Vibra tion S umm ary

MAXIMUM RA_&E 19HERTZ RANGE 45 HERTZ RANGE 15 HERTZ RANBEMEAS. LEVEL TIME LEVEL TIME LEVEL TIME LEVEL TIMENO. AREA FDNITORED GPEAK (SEC) GPEAK (SEC) GPEAK (SEC) GPEAK (SEE)

E091 Fwd FieldSplice- Thrust 1.20 2 0,08 620 0.10 9483 O.Ol 9550

E099 Fwd BendingMode - Pitch 0.58 9483 0.04 610 0.58 9483 0.01 9550

E]O0 FwdBendingMode - Yaw O.B2 9483 0.07 600 0.52 9483 0.01 9550

EBB2 AftSeparationlane- 0.58 87 0.16 620 0.08 9483 0.04 9550T h r u s t

A010 GimbaI Block - Pitch 0.06 6 0.01 620 ... .. .. .. .. .. ... .

iAOII GimbalBlock- Yaw 0.14 6 0.01 620 0.01 9483 0.03 9550

A012 GimbalBlock- Thrust 0.30 620 0.30 620 0.08 9483 O.O6 9550

E251 J-2 ChamberDome- Thrust 0,82 9220 0.25 620 0.05 9520 0.07 9650

AOI3 J-2Enginekirt-Pitch 0,21 568 0.17 620 0.10 9483 0.13 9550

AOI4 J-2 EngineSkirt- Yaw 0,25 56B 0.21 620 0.13 9483 0.17 9550

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 pre-

dictions. (See Figure 9-9.)

9.3 VIBRATION EVALUATION

9.3.1 S-IC Stage and Engine Evaluation

Structure, engine, and component vibration measurements taken on the S-IC

stage are summarized in Table 9-2 and in Figures 9-I0 through 9-12. A

total of 44 single sideband vibration measurements were recorded, of which

42 yielded valid data throughout flight. Measurement locations are shown

in Figure 9-13.

9.3.1.1 S-IC Stage Structure. Stage structure vibration data exhibited

composite RMS levels and spe(_ra shapes within the data envelopes of pre-

vious flights. The AS-505 maximum inflight intertank and forward skirt

structure RMS levels lag those measured on previous flights because the

Max Q region occurred later in flight.

9.3.1.2 F-l Engines. The F-l engine combustion chamber and turbopump

measurements compare closely with previous flight data in both overall

levels and spectra shapes. Measurement E038-101 shows a high Grms level

when compared to previous valid data; consequently, it is questionable

and is not included in the engine turbopump plot.

9.3.1.3 S-IC Components. All S-IC component vibration measurements were

valid, and the levels measured agreed with those measured on previous flights.

9-11



L EG EN D

ANALYSISMODE 1 --

A N A L Y S IS M O D E 2

lO ANALYSISODE3--

COMMAND MODULE ANALYSIS MODE _4.-ME AS UR ED MO DE ®MODE NUMBER N

8

,.

_ J_ ,.t

_4"- f_ ® ._ ®

_ - ®_ @- --_2_"........ <7)

0

0 20 40 60 80 lO0 120 140 160

R ANGE T IME, SECON DS

Figure9-9. AS-505LateralAnalysis/MeasuredModal FrequencyCorrelation

9.3.2 S-II Stage and Engine Evaluation

Comparisonsof Grms valuesfor AS-505and previousflightdata are shown

on Table 9-3 and in Figures 9-14 through 9-16. The AS-505 peak level

at liftoff for the interstage frames radial measurement and all AS-505

values shown for the aft skirt stringers radial vibration measurement

were determined from contractor digitized data. All other values shown

were determined from NASA Grms history data. The variations between

the five flights are considered normal.

9.3.2.1 S-If Stage Structure. In general, the S-II stage structure

vibration levels were within the envelopes established by previous

flights, The forward skirt stringers tangential vibration and aft skirt

stringers radial peak value vibration levels were slightly above the

envelopes established by previous flights. The interstage frames

tangential vibration data were invalid and are not included in

Figure 9-14.

9.3.2.2 S-If Stage J-2 Engines. The S-II engine combustion domes longi-

tudinal and L H2 pumps radial vibration envelopes (F igure 9-15) show a

9 - 1 2



T able 9 -2. S -IC S tage V ibration S um m ary

MAXIMUMGRMSAT RANGETIMEOVERAL L

PREVIOUS FLIGIIT G_S

MEASUREMENT DATA AS-505 LIMIT REMARKS

STRUCTURE

Thrust StructureE023-115 14.7 at 0 10.8 at O 22

E024-115 11.2 at O 14.1 at -2.0 25

E053-I15 6.9 at 149.5 5.5 at 156.0 17

E054-115 3.7 at 150 17 AS-505 data are invalid.E079-I15 3.3 at 148 3.1 at 158.0 17

E080-115 4.2 at 148 3.7 at 158.0 17

IntertankStructure

E020-118 7.7 at 2 5.6 at -2.0 87

E021-118 9.1 at 4 9.4 at 0 27

F o rw ar d Ski rt

Structure

E046-120 3.6 at 94 5.0 at 85 30 Located near c(_nmand

E047-120 6.1 at 3.9 5.2 at 5.7 30 destruct vibrationisolatedpanel.

ENGINE

CombustionChamberE036-I01 8.8 at 20.5 7.58 at 156 49

E036-102 9.7 at 0 8.01 at 110.2 49E036-I03 8.3 at 53 8.38 at 10.2 49E036-I04 8.4 at 106.8 7.22 at 120.3 49E036-I05 8.2 at 130.5 8.03 at 50.3 49

TurbopumpE037-I01 41.5 at20.0 23.8 at 157.0 41

E038-I01 39.0 at 1.0 41 Data questionable,due to anamplifier calibrationerror.

E039-I01 26.5 at 125.0 16.8 at 132.4 41

E040-I01 17.3 at 123.8 15.4 at 154.0 41 Data containsspikes ata11 analysis times.

E041-101 20.9 at 158.0 17.6 at 157.0 41E041-I02 17.5 at 144.5 18.8 at 154.0 41

E042-I02 9.6 at 86 8.6 at 157.0 41E042-I03 10.9 at 148.1 9.I at 154.0 41_ Data containsspikes at

E042-I04 11.2 at 79.0 9.4 at 144.5 41J all analysis times.042-I05 10.7 at 26.6 9.0 at 126.2 41

COMPONENTS

EngineActuatorsE030-I01 9.4 at I11 4.4 at 125.5 30E030-I02 5.0 at 123 4.6 at O 30

E031-I01 6.7 at 118 6.0 at 156.0 30

E031-I02 7.8 at 107 6.7 at 134.3 30

E032-101 15.1 at 111 7.4 at 15O.O 30

E032-I02 14.0 at 89 13.8 at 155.0 30

E033-101 8.8 at 100 8.5 at 71.4 30

E033-102 7.0 at 127 6.3 at 15g 30EO34-101 5.3 at 124 4.1 at 150 30

E034-102 5.5 at 135 10.2 at 0 30

E035-101 15.0 at 68 14.7 at 71.3 30

E035-I02 10.5 at 127 8.4 at 134.3 30

lleatShield Panels

E1O5-10(i 76.6 at -i 73.2 at O 33

E106-I06 70.8 at O 70.9 at 0 33

E107-106 74.4 at 0 70.2 at 0 33

PropellantDeliverySystem

E025-I18 2.7 at 132 1.7 at -2.0 9

E026-I18 3.1 at 118 2,8 at 0 9E027-115 10.4 at -0.5 8.3 at -2.0 22E028-115 11.3 at 118 10.D at 0 22

9-13



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PREVIOUSFLIGHT _ AS-505 FLIGHTATAENVELOPE DATAENVELOPE

60- ENGINETURBOPUMP

[I.

_NNNNNNNNN _N_ _N _

15-

c_ ENGI'NEOMBUSTIONCHAMBER

I I I

--I S ....

¢_1

20 40 60 80 100 120 140 160

RANGETIME, SECONDS

F igure 9-11 S -IC Stage Engine Vibration Envelopes





E037-101 E039-10I

ZE038-I0] E040-101 E041-10]

..... --STA I54] E036-lOI E041-I02

i _ "_ E036-102

E046-120 EO36-IOS_E042-102

E036-104 E042-103

FORWARD SKIRT -- STA 1401 E042-I05

E064-120 O __ --

E063-120 I

POS I

E034-101

EO_-I02 EO30-IOZE030-101 E035-IOI

LOX T_K E03S-lO2 E032-102

E093-119_._ E032-1 1

--STA 912 E031-101-- E031-102E033-102

EO25ZIl_ - . F-I ENGINE

E026-II FIN C

R02l-lt8 POS _ IPOS Il l

EIOS-106

,il ----STA g12

FIN D-- FIN B

[ii INTERT_K E059-118

DIOS-i.POS I' "PO_ 1]

--STA 602 I "EI06-1O6FIN A

EODI-I1H HEAT SHIELD062-118

FIN D

E086-117-- EO_-117 -- STAT]_ 2_E087-117

S_ 602-

E091-115

ii-i _ " i _'8--11/--STATI_ 112

FIN C-- --- FI " EO82-I]_o?

THRUSTSTR_TUREPOS IV / "_POS I E014-115

E085- ]17

FJEL T_K E023-115

EO13-IIS

1!024-115 STATION280

EOG7-IIS

E055-115

Ill . E056-115-II-STATI_ 112

E054-]15

E079_115

E053-115 _EOBO-IISIIV

Figure 9-13. S-IC Vibration Measurement Locations

9-17



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1_7 MACH1 _ S-I[ ESC _ pHR5H|FT_7 MAXQ 5-[I CECO S-I| OECO

"AS-5QS .... PREVIOUSLIGHTS

I,_ FOR_ARC'KIRTSTRINGERS,RAOtAL 141 ENGINEI BEAH,LATERAL

- >4 _'

! "/>_l'1 /i s- ....

FOR_4ARDKIRTSTRINGERS,TANGENTIAL

._. _._ THRUSTCONELO_GERONS,ONGITUDINAL .__ .--

_ I _ _ -'-F--4---4----_---t---,L--!I6 ......

>_---4 FOR]_ARD"_ SKIRT/,*,ISTRI_GERS,I_NGITUOI/e,LI

_ ;,1 -:....._......._....AFTSKIRTSTRINGERS,RADIAL 18

Io I_ .

_ g

z

.. ,'? , _- ._

,,i:, __...................,,."' "--_,... I,g_

-so 0 50 100 150 2oo 2so 3oo 350 40o 450 500 550 _50 0 so 100 150 200 zso 3oo 350 4oo 450 500 _seRANGETIME, SEC0f_0S p,NGETIME, SECONDS

Figure 9-14. S-II Stage Structure Vibration Envelopes

9 - 1 9





MACH _ S-IfESC _ PHRSHIFT_X Q S-II CECO S-II OECO

AS-SOS .... PREVIOUSLIGHTS

FOf_AROSKIRTCONTAIFYEREOIRPUT, _ORMAL FORWARDKI RTCONTA_t_RZZI II_PUT,NORt'_LB

c_ _ I0

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81 FORWARDSKIRTCONTAINER220RESPONSE,NORMAL

_ _4 g4

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" FORWARDSKIRTCONTAINER223 RESPONSE,NORMAL RI

oSO 0 50 IO0 150 200 E50 300 350 400 ASO 500 SSO "50 0 50 100 lEO 200 250 300 350 400 _50 500 550

RANG_TIME, SECONDS P,NP._TIME, SECONDS

F igure 9-16. S-II Stage Component Vibration Envelopes

(SheetI of 2)



HACH1 _ $-II ESC _ PNRSHIFTlAX 0 S-If CECO $-II OECO

-- AS-S05 .... PflGVIOUS FLIGHTS

_-o k I _i I I

_ _ _ k _-___- •.............

DTHou..,_ETINE.OO.ON?OT_ _ h _--'_ THRUSTCONECONTAINER207 RE PONSE, NORMAL

_ ,_ ........... ......... .,.... _ _.... ,_;-_.......... _-:..=.-..... _-- ....

_. T_,D_,oNiNTi'"'_I'"_,o''.__,,o.oNEON,"rNER0'ES_'ON,.ONG'_O'N"i'. s_ ,\"::' I $_d ;I _ ,- -4- ..... . ".. ... ... ... ..

o.¢

_ _ LH2 I_EVPJ.VE, LONGITUDINALO

_' TH._uE..o..._o,,',',.NE,.,REo.s_ .. _ L.2:',_EV_,'E.O_.,L

o:_;_ I I__. _ .... .":',, -_; :_._' _ ,_....... .o_ _ .. ..-,,.____o_ _

_¢'_ THRUSTcoNE CONTAINER206 RESPONSE,LONGITODIrlAL ! LOXSL_P I_REVAL_tE,RAbIAL

•,_ I , -_ _ , , =- ..'_...k,x..... _ ......... _"'" .... ""- _ 7o-EO O 50 100 150 200 250 300 350 400 450 500 550 o < -EO 0 50 100 150 200 250 300 350 400 450 500 55

RANGETIME, SECONDS RANGETZHE, SECONDS

Figure9-16. S-II StageComponentVibrationEnvelopes(Sheet2 of 2)



drop in vibration level at CE CO . A sim ila r drop in level fo r the engineN o. 5 L O X pum p is not included in the figure because the reduced levelremained within the envelope of the other four engines.

9 .3.2.3 S -II S tage Components. In genera l, all S -II stage componentvibration 'levels, as shown in F igure 9-1 6 , agreed closely with the

previous flight data.

9.3.3 S-IVB Stage and Engine Evaluation

T wo vibration measurements were made on the structure, 15 at components

mounted on the stage, and 13 at engine components. The maximum composite

levels are indicated in Figures 9-17 and 9-18 and in Table 9-4.

9.3.3.1 S-IVB Stage Structure and Components. The envelope of vibra-

tion levels for the stage structure and components is shown in Figure 9-17.

The data of the figure show the range of vibration levels at the input to

components mounted on the forward and aft sections of the stage. TheA S-505 levels were lower than the maximum measured during the A S-503 flight.

9.3.3.2 S-IVB Stage J-2 Engine. Data measuredduringthe AS-505 flighton the two turbopumps and the combustion chamber dome are shown in Figure

9-18. The AS-503 levels presented for comparison include data from the

turbopumps only. The differences between the measured vibration environ-

ment are within the normal scatter of the engines.

AS-505 S-IVB first and second burn data from components on the J-2 engine

are shown in Figure 9-18. In addition, the nominal range of levels from

similar measurements monitored on the AS-503 flight are shown. The AS-505

levels are within the range of the AS-503 data.

9.3.3.3 S-IVB Stage ASI Lines Dynamics. Dynamic strain measurements

were made on the LOX and t.H ASI lines. The LOX ASI line strains ranged

from 9 to 34 win/in.RMS (AS-503flight linestrainsrangedfrom lO to 20

uin/in.RMS). The LH2 ASI line strainsranged from 17 to 60 win/in.RMS(AS-503flight linestrainsrangedfrom 20 to 50 win/in.RMS).

9-23



F O RW A RDC O M PO N E N TSA N D S T RU CT UR E (F IE L D S P LIC E)

I0 I I 1FIRSTBURN SECONDURN

_ AS-503 K-_ AS-505o AFT COMPONENTSNDSTRUCTUREAF SKIRT_ 15

I

_ 10

5

\_ i\ _\_\ II I// / / //

0O 40----80 12O 500 600 700 9200 9300 )500 9600

_N GE T IME , S EC ON DS

' ' ,_ ' _ Io_ '00:02:00 00:10:00 02:350 02:40:00

N GE T IME , HO UR S: MIN UT ES :S EC ON DS

Figure 9-17. S-IVB Stage Vibration Envelopes



-2 E NGIN E

60 FIRSTURN SECONDURN

5o [

20 11/,'III, III

. I0 _"- AS-505CHAMBEROMENODATAONAS-503): r a -

G. ' "V '

_-- _ AS-503 _ AS-505< J -2 E N G IN E C O M P O N E N T Sr_uJ 6 0

50 MEASUREMENTSOT_ MONITOREDURING

Y/P2,............0 S- I C BURN .,"".".",.".'." ."" . . . _ ////////// /,/////// ////

30 ' '_ "3<%_V_ZvvO<)<]

20 LEVELS ___: ____

INSIGNIFICANT __._: ,_ _ _.,_"I0 DURINGS-II BURN

o ",,550 600 650 700 9200 9300 9400 9500 9600

R AN GE TIM E, S ECO NDS

I I I 121:0/_'0" I I I 1 I I I I:lO:O0 O: 2:34:00 2:36:00 2:38:00 2:40:00

R A NG E T IM E . H O UR S :M IN UT ES :S E CO N DS

F igure 9 -1 8. S - IVB S tage E ngine V ibration E nvelopes



T able 9 -4. S -IVB Vibration S ummary

MAX RANGE

AREAMONITORED LEVEL TIME REMARKS

GRMS (SEC)

Structure Field Splice at Position I 0.8 1.5 Frequency Limited to 220 Hz

( L ow F re qu en cy ) , T hr us t

Station 69.9 m (2748 in.) at 0.4 87 Frequency Limited to 220 Hz

Position II on A ft S kirt, T hrust

Engine Combustion Chamber Dome 0.6 9220 Frequency Limited to 220 Hz

( L ow F r eq ue nc y) , T hr us t

Combustio_ Chamber Dome 9.4 560

Longitudinal

LH2Turbopump,Lateral 25 9260

LOXTurbopump,Lateral 50 9210

Stage InputtoLH2 VentDisconnect, 2.5 85

Components Fwd Skirt, Thrust

(Forward)Input to LH2 Vent Disconnect, 3.5 84Fwd Skirt,Radial

InputtoContinuousent 3.6 0.5

Module, Fwd Skirt, Radial

Stage HeliumBottle,Thrust 3.4 700

Components Structure, Pitch

(Aft)

Input to LH2 Feedline at 2.0 9216

LH2 Tank, Thrust

Input to LH2 Feedline at 2.9 67

LH2 Tank, Radial

Input to LH2 Prevalve in LH2 2.6 9219F ee dl in e, T hr us t

Input to LH2 Prevalve in LH2 2.1 560Feedline, R_dial

AmbientPanel,Inputto 1.4 0.5

C hi ll do wn I nv er te r, T hr us t

AmbientPanel,Inputto 4.0 0.5

C hil ldo wn In ve rt er , R adi al

APS, Input to Propellant 4.6 83C on tr ol Mo du le , R adi al

APS, InputtoPropellant 5.7 83Control Module, Tangential

APS, Input to Helium 9.6 96

R eg ulator, Tangential

Input to RetrorocketFwd 3.2 0.5

S uppor t, A ft

Input to LOX ChiIldown Pump, 3.9 0.5Aft LOX Dome, Normal to Dome

9 - 2 6



Table 9-4. S-IVB Vibration Summary (Continued)

MAX RANGE

AREAMONITORED LEVEL TIME REMARKS

GRMS (SEC)

Component Main Fuel Valve, Tangential 6.5 702

J -2 E ng in e Main Fuel Valve, Radial 6.5 702

Main Fuel Valve, Longitudinal 11.5 702

LOXTurbineBypassValve, 8.4 506 MeasurementFailedSecondBurn

Tangential

LOX TurbineBypassValve, 8.6 560Radial

LOX TurbineBypassValve, 18.2 702

Longitudinal

ASILOXValve,Radial 14.7 702

ASILOX Valve,Longitudinal 21.6 702

FuelASIBlock,Radial 36.7 560

9-27/9-28



S E CT IO N I0

G UIDA N CE A N D N A VIG AT IO N

I0 .I S U M M A R Y

1 0 .1 .1 F ligh t P rogram

The guidance and navigation system performed satisfactorily during allperiods for which data are ava ilable. T he boost nav iga tion and gu idance

schemes were properly executed, and translunar trajectory in jection para-

meters were within tolerances. All orbital operations were nominal andS -IV B stage safing was satisfactorily accom plished, resu lting in a helio-

centric orbit for the S-IVB/IU as planned.

1 0 .1 .2 Instrum ent Unit C om ponents

T he L aunch Veh icle D ig ita l C om puter (L V DC ), the L aunch Veh icle Data A dapter

(L VDA ), and the S T -1 24M -3 inertia l p latform functioned sa tisfacto rily. N oanomalies or deviations have been discovered.

1 0 .2 G UID A N CE C O M PA R IS O N S

T he postfligh t gu idance hardware error ana lysis was based on comparisons

of the S T -1 24M -3 p latfo rm m easured velocities with the observed postfligh t

tra jectory established from externa l tracking data . N o precision tracking

data were ava ilable and the boost-to -parking orbit trajectory was established

by a com posite fit o f C -band radar da ta . F igure I0 - I p resen ts the com pari-

sons of the p la tfo rm m easured velocities with correspon ding va lues from the

fina l observed postfligh t tra jectory. A positive difference indicates tra-

jectory data greater than the p la tfo rm m easurem ent. A ltho ugh the overa lldifferences are relatively small, they do not reflect a characteristic trend

fo r p la tfo rm hardware errors. T he differences probably reflect m ore tra -jecto ry erro r than gu idance erro r. T he velocity differences a t S -IV B first

Engine Cutoff (ECO) were -0.13 m/s (-2.5 ft/s), 1.0 m/s (3.3 ft/s), and -0.6 m/s(-2.0 ft/s) for altitude, crossrange, and downrange velocity, respectively.

Due to lim ited track ing coverage of the second burn m ode, tha t portion o f

the observed postfligh t tra jectory was constructed by in itia lizing the sta tevector and integra ting the platfo rm -m easured velocities. A ny velocitydifferences for the second burn were due to data transfo rm ation and inter-

polations.

1 0 - 1



_7 S-ICOECO %_7S-IIOECOR_7S-If CECO _S-IVB FIRSTECO

-2 -6 _ ='

2

_ -I --3 <,-,_

_ -2i --6_-_

E -6

NN-I --3_."_

O lO0 200 300 400 SO0 600 700

RANGETIME,SECONDS

Figure lO-l. Tracking and ST-124M-3 Platform Velocity Comparison

(Trajectory Minus Guidance)

Velocities measured by the ST-124M-3 platform system at significant flight

event times are shown in Table lO-l, along with corresponding values com-

puted from the final AS-505 observed postflight trajectory and the preflight

operational trajectory. Since the same thrust profile was used in the pre-

flight and postflight operational trajectories, the inertial platform out-

puts should be equivalent. The differences between the telemetered velo-cities and the observed postflight trajectory values reflect some combina-

tion of small guidance hardware errors, tracking errors, and errors in

interpolating data for event times. Thejifferences between the telemetered

and operational trajectory values reflected off-nominal flight conditions

and vehicle performance.

C omparisons of navigation (PA CS S 13 coordinate system) positions, velocities,

and flight path angle at significant flight event times are presented in

T able I0-2. The guidance (L VDC ) and observed postflight trajectory values

are in relatively good agreement throughout the flight. The component

differences at parking orbit insertion and at Translunar Injection (TLI)

are given in Table I0-3.

10-2



Table I0-I. Inertial Platform Velocity Comparisons

VELOCITYM/S (FT/S)**EVENTS DATA

S OUR C E A L TITUDE (Xm) CR O S S R AN GE (Y m)DOWN R A N GE (Zm)-

Guidance 2571.64 7.90 2246.55S-IC (8437.14) (25.92) (7370.57)

OECO Postflight 2572.05 7.98 2245.74

Trajectory (8438.48) (26.18) (7367.91)

Preflight 2583.50 -0.54 2226.35Trajectory (8476.05) (-I.77) (7304.29)

Guidance 3470.47 -3.60 6751.83

S-II (II,386.06) (-ll.Sl) (22,151.67)

OECO Postflight 3470.27 -3.73 6751.08

Trajectory (II,385.40) (-12.24) (22,149.21)

Preflight 3462.06 -0.72 6763.89

Trajectory (II,358.46) (-2.36) (22,191.24)

Guidance 3210.19 2.05 7611.70

(10,532.12) (6.73) (24,972.77}

First Postflight 3209.43 3.06 76li.08

S-IVBECO Trajectory (10,529.63) (I0.04) (24,970.73)

Preflight 3206.03 1.54 7610.76

Trajectory (I0,518.47) (5.05) (24,969.68)

Guidance 3209.50 2.05 7613.35

(I0,529.86) (6.73) (24,978.18)

ParkingOrbit Postflight 3208.78 3.10 7612.55

Insertion Trajectory (I0,527.49) (lO.17) (24,975.56)

Preflight 3205.37 1.55 7612.46

Trajectory (I0,516.30) (5.08) (24,975.26)

Guidance 3079.06 204.90 -696.22

(lO,lOl.90) (672.24) (-2284.19)

Second :Postflight 3078.96 204.28 -695.95

S-IVBECO* iTrajectory (I0,I01.57) (670.21) (-2283.30)

iPreflight 3094.33 14.53 -628.81

Trajectory (I0,152.00) (47.67) (-2063.02)i

• Second burn velocity data represent accumulated velocities

from Time Base 6

• * PACSS 12 Coordinate System

NOTE: Preflight trajectory data were adjusted for trajectory

error in platform values at liftoff.

I0-3



Table I0-I. Inertial Platform Velocity Comparisons (Continued)

EVENTS DATA VELOCITYM/S (FT/S)**

SOURCE ALTITUDE (Xm) CROSSRANGE(Ym)DOWNRANGE (Zm)

Guidance 3083.00 205.25 -696.60

(10,114.83) (673.39) (-2285.43)Translunar

Injection * Postflight 3083.01 205.35 -696.53Trajectory (10,114.86) (673.72) (-2285.20)

Preflight 3098.17 14.51 -629.10

Trajectory (I0,164.60) (47.60) (-2063.98)

* S econd burn velocity data represent accumulated velocitiesfrom Time Base 6

** PACSS 12 Coordinate System

NO TE : Preflight trajectory data were adjusted for trajectory

error in platform values at liftoff.

I0-4



T able 1 0 -2. Guidance Comparisons

POSITIONS VELOCITIES FLIGHTATH

METERS M/S ANGLE (DEG)DATA (FT) (FT/S)EVENT

SOURCE

Xs Ys ZS R XS is ZS Vs

Guidance 6,436,333 39,704 158,973 I 6,438,417 827.45 128.50 2621.08 2751.72 18.930

(21,116,578) (130,261) (521,564) [21,123,416)(2714.72) (421.59) (8699.35) (9027.96)

S-IC Observed

OECO Postflight 6,436,404 39,621 158,904 6,438,488 828.25 128.58 2621.16 2751.91 18.9457

Trajectory (21,I16,812) (129,990) (521,337) (21,123,647) (2717.36) (421.85) (8599.61) (9028.58)

Operational 6,436,982 39,195 156,355 6,439,000 863.38 120.16 2602.10 2741.10 19.5450

Trajectory _(21,I18,708) (128,593) (512,977) (21,125,328) (2799.80) (394.23) (8537.07) (8993.11)

Guidance 6,283,497 81,690 1,880,904 6,559,482 -1893.09 86.98 6633.39 6898.82 0.74600

[20,615,150)268,010) 6,170,946)(21,520,610)-6210.92) (285.38) (21,763.08)(22,633.92)m

cm S-II Observed

OECO Postflight 6,283,760 81,586 1,880,627 6,559,65_ -1893.09 86.90 6632.82 6898.24 0.74107

Trajectory 120,616,009) (267,669) (6,170,037) (21,521,172) (-6210.93) (285.10) (21,761.22) (22,632.02)

Postflight

Operational 6,281,468 81,521 1,891,355 6,560,542 -1908.57 90.21 6642.10 69|1.46 0.7346

Trajectory (20,608,492) (267,458) (6,205,231) (21,524,088) (-6261.71) (295.97) (21,791.67) (22,675.40)

Guidance 5,882,772 93,925 2,908,961 6,563,374 -3454.27 75.80 6983.64 7791.60 0.00179

(19,300,431) (308,153) (9,543,835)(21,633,379)(-II,332.91) (248.69) (22,912.20) (25,562.99)

First Observed

S-IVBECO Postflight 5,882,950 93,881 2,908,740 6,563,435 -3464.92 76.89 6983.19 7791.42 -0.0064

Trajectory (19,301,018) (308,010) (9,543,109) (21,533,581)(-11,335.04) (252.26) (22,910.73) (25,562.40)

Postflight

Operational 5,883,463 93,919 2,907,773 6,563,467 -3452.94 75.83 6984.02 7791.35-0.00020

Trajectory (19,302,701) (308,132) (9,539,936) (21,533,686)(-II,328.54) (248.78) (22,913.46) (25,562.17)



T abl e 10-2. Gu ida nce C om pa ris ons ( Co nti nue d)

POSITIONS VELOCITIES FLIGHTATH

METERS M/S ANGLE (DEGIEVENT DATA (FT) (FT/S)

SOURCE

Xs Ys Zs R Xs Ys Zs Vs

5,84_,815 94,677 2,978,600 6,563,38(I -3537.6( 74.60 6943.82 7793.41 0.00260Guidance (19,185,745) (310,619) (9,772,308) (21,533,400 :(-11,606.50)i(244.75) (22,781.56) (25,558.93)

Parking Observed

Orbit Postflight 5,847,847 94,642 2,978,439 6,563j33! -3538.19 75.72 6943.18 7793.09 -0.00494

Insertion Trajectory 19,185,849) (310,506) (9,771,781) (21,533,252 i-II,608.23)(248.43) (22,779.46) (25,567.88)

Postflight

Operational 5,848,514 i 94,671 2,977,423 6,563,46! -3536.30 74.63 6944.23 7793.16 O.OOlOTrajectory (19,188,037)! (310,600) (9,768,447) (21,533,690 _-11,602.03) (244.85) _22,782.91) (25,568.11)

Guidance 191,6241 -123,594 -6,690,743 6,694,620 10,799.27 230.74 -I003.50i ]0,848.24 6.92400

(628,688) (-405,493) -21,951,2701 (21,963,976 (35,430.67) (757.02) (-3292.32)(35,591.34)i Second Observed

S-IVB Postflight 188,2031 -]24,409 -6,692,91_ 6,696,719 10,797.04 229.34 -]DO9.00 10,846_56 6.927

ECO Trajectory (617,462)(-408,167) -21,958,39]) (21,970,863 (35,423.76)(752.43) (-3310.37) (35,585.83)

Postflight

Operational 172,344! -123,812 -6,690,55( 6,693,920 10,798.20 231.87 -1024.03 I0,849.12 6.8673

Trajectory (565,432)(-406,208) -21,950,643) (21,961,68l (35,427.16) (760.72) (-335g.68) (35,594.23)

Guidance 299,6131 -121,275 -6,700,34( 6,710,259 10,799.97 232.66 -915.13 I0,841.17 7.37843

(982,982) (-397,885) -21,982,743).(22,015,286 (35,432.97) (763.32) (-3002.40)(35,568.14)

T ran_lun ar O bserved

Injection Postflight 296,191 -]22,100 -6,702,54_ 6,710,200 ]0,797.9I 23].99 -920.85 10,839.59 7.379Trajectory (971,755)(-400,589) -21,989,987)(22,015,093 (35,426.21)(761.12) (-3021.16)(35,562.96)

Postflight

Operational 280,3491 -121,482 -6,700,355 6,707,318 ]0,799.01 233.85 -935.57 10,841.98 7.3219

Trajectory (919,780) (-398,562) 1-21,982,793_(22,005,637 (35,429.83) (767.23) (-3069.45)(35,570.80)



Table 10 -3. Guidance Component Comparisons

P A R A M E T E R S O B S E R V E D -G U ID A N C E P O S T F L IG H T -G U ID A N C E

P AR KIN G O R BIT IN SE RTIO N DIFF ER EN CE S

Axs m/s (ft/s) -0.53 (-I.73) 1.36 (4.47)

Ays m/s (ft/s) 1.12 (3.67) 0.03 (0.I0)

Azs m/s (ft/s) -0.64 (-2.10) 0.41 (1.35)

AVs m/s (ft/s) -0.31 (-I.02) -0.25 (-0.82)

ARm(ft) 45.0(148.0) 89.0(290.0)

Aedeg -0.00754 -0.0016

T RA N S LU NA R IN JE CT IO N DIF FE RE N CE S

Axs m/s (ft/s) --2.06 (-6.76) -0.96 (-3.14)

Ays m/s (ft/s) -0,67 (-2.20) 1.19 (3.91)

A_s m/s (ft/s) -6.72 (-18.76) -20.44 (-67,05)

AVs m/s (ft/s) --I,58 (-5.18) 0,81 (2.67)

_R m (ft) -59.0 (-193.0) -2941.0 (-9649.0)

Aedeg 0.00277 -0.0546

The ST-124M-3 platform measurements and the LVDCflight programs werehighly successful in guiding the A S -50 5 vehicle to near nom inal end

conditions. A m inimum of corrections were required for the spacecraft toaccomplish its mission,

1 0.3 N AVIGA TXO N A ND GUIDA NCE S CHE M E E VA LUA TIO N

A ll analyzed guidance performance measurements indicated satisfactoryguidance during S -IVB first and second burns. T he active guidance phasesstart and stop times are given in Table 10-4, Included in this table arethe start and stop tim es for the artificial tau phases and chi freezes.T he minor loop chi attitude commands and orbital guidance comm ands aregiven in F igures 1 0 -2 and 1 0 -3, respectively. T he lower than predictedgeocentric radius and the higher crossrange velocity component at IterativeGuidance M ode (IGM ) initia tion were com pensated for by com m anding pitch

approximately 2 degrees more positive than predicted and yaw approxi-m ately 0 .5 degree m ore negative than predicted, as shown in F igure 1 0 -2.T he deletion of the attitude freeze resulted in a sm oother transition

from S -II to S -IVB guidance phases than the sam e transition during

A S -503 and A S -504 missions. P re-IGM guidance functioned satisfactorilyas programed. O rbital guidance events for which telemetry was avail-able were _ccomplished satisfactorily. The guidance during S-IVB secondburn resulted in satisfactory T L I parameters as shown in T able 1 0 -5,

1 0 - 7



T able 1 0 -4. S ta rt and S top T im es fo r IG M G uidance C om m ands

STEERING

M IS A L IG N M E N T T E R M IN A LEVENT* IGMPHASE ARTIFICIALAU CHIFREEZE

CORRECTION GUIDANCE

(SEC) (SEC) (SEC) (SEC) (SEC)

START STOP START STOP START STOP START STOP START STOP

FirstPhaseIGM 202.9 484.8 222.6 491.9

SecondPhaseIGM 484.8 552,7 484.8 490.2 493.8 552.7

ThirdPhaseIGM 552.7 695.7 560.I 568.9 567.1 695.7 669.4 697.3 697.3 704.9**

FourthPhaseIGM 9218.2 19333.2 9223.9

Fifth PhaseIGM 9333.2 9547.7 9333.2 9339.6 9549.2 9521.1 9549.3 9549.3 9568.7**

* A|I times are for the start of the computation cycle in which the event occurred.

** Start orbital time line.

Control parameters indicate slight attitude perturbations at IGM initia-tion and S-IVB Programed Mixture Ratio (PMR) shift. The perturbationswere expected and were not significant. The minor loop satisfactorilyconverted the guidance commands into steering signals throughout themission.

I0.4 GUIDANCESYSTEMCOMPONENTVALUATION

lO.4.1 LVDC Performance

The LVDC performed as predicted for the AS-505 mission. No valid errormonitor words and no self-test error data have been observed that indi-

cate any deviation from correct operation.

I0.4.2 LVDA Performance

The LVDA performance was nominal. No valid error monitor words and noself-test error data indicating deviations from correct performance wereobserved.

I0.4.3 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.

I0-8



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-60 _ 0 YAWATTITUDEOMMAND..... OPERATIONALTRAJECTORYPREDICTED

-80 D PITCHTTITUDEOMMAND.... OPERATIONALTRAOECTORYPREDICTED

-IOO A ROLL ATTITUDE COIV_tAND

-__ OPERATIONALTRAaEETORYPREDICTED

-120 _ OPERATIONALRAJECTORYREDICTIONS&

\ AVAILABLETO 11,5(_ SECONDS

-140 \160 %%

-]8C0 2 4 6 8 10 12 14 16 18 20 22 24 26 28 30

RN_GE TIME,SECONDS

L | I I I I I I. I I I0 0:50:00 1:40:00 2:30:00 3:20:00 4:10!00 S:O0:O0 5:50:00 6:40:00 7:30:00 8:20:00

RANGETIME,HOURS:MINUPES:SECONOS

Figure 10-3. Orbital Attitude Commands



Table 10-5. Translunar Injection Parameters

POSTFLIGHT TRAJECTORY LVDCMINUSMINUS LVDC PREDICTED

PARAMETER PREDICTED TRAJECTORY PREDICTED

I nertial Velocitym/s I0,84].98 10,839.59 -2.39 10,841.17 -0.81

(ft/s) (35,570.80) (35,562.96) (-7.84) (35,568.14) (-2.66)

F li ght P at h A ngl e

deg 7.322 7.379 0.057 7.378 0.056

D es ce ndi ng M ode

deg 123.537 123.515 -0.022 123.527 -O.OlO

Inclination

deg 31.691 31.698 0.007 31.698 0.007

Eccentricity 0.97836 0.97834 -0.00002 0.97830 -0.00006

C3m2/s2 -I,307,603 -I,308,471 -868 -1,310,867 -3,264

(ft2/s2) (-14,074,922) (-14,084,267) (-9345) (-14,110,055) (-35,133)

1 0 .4.4 T elem etry O utputs

A na lys is o f the ava ilable L V D A te lem etry buffer an d fligh t contro l co m -

puter attitude error p lo ts indica ted sym metry between the buffer outputsand the ladder ou tputs. T he ava ilable L V DC power supply plots indicated

satisfactory power supp ly performance. T he H6 0 -6 0 3 gu idance com pute r

telemetry was completely satisfactory.

I0.4.5 Discrete Outputs

No valid discrete output register words (tags 043 and 052) were observed

to indicate guidance or simultaneous memory failure.

I0.4.6 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 R edundant (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.

10.4.7 ST-124M-3 Inertial Platform Performance

The inertial platform system performed as designed. The inertial gimbal

temperature fell below specifications; however, there are no indications

lO-ll



of degraded inertia l p latform performance. T he temperature went belowthe minimum specification of 31 3.1 5°K (I0 4.0 °F ) at I0 ,0 0 0 seconds,reaching 310 .1 5°K (9 8.6°F) at approximately 25,0 00 seconds.

The accelerometer servo loops functioned as designed and maintained the

accelerometer float within the measuring head stops (±6 degrees) through-

out the flight. T he accelerometer encoder outputs indicated that theaccelerometers accurately measured the vehicle acceleration.

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 any time during the mission.

1 0 - 1 2



S E CTIO N II

C O N TR O L S Y S T EM

II.I S UM M A R Y

T he A S -50 5 F light Contro l Com puter (F C C), T hrust Vector Contro l (T VC),and A uxiliary P ropulsion S ystem (A P S ) satisfied all requirements forvehicle attitude control during the flight. B ending and slosh dynam ics

were adequately stabilized. The preprogrammed S -IC boost phase yaw,roll, and pitch maneuvers were properly executed. The S-IC outboard

cant was accomplished as planned.

T he peak winds observed during the flight were slightly less than the95-percentile M ay wind and were well within the capabilities of the con-tro l system. T he m axim um pitch and yaw engine deflections were causedby wind shears.

S-IC/S-II first and second plane separations were accomplished with nosignificant attitude deviations. A t Iterative Guidance M ode (IGM ) initi-ation a pitch up transient occurred sim ilar to that seen on previousflights. A t S -II early Center E ngine Cutoff (CE C O ), the guidance param -eters were m odified by the loss in thrust. T here was a change in yawattitude due to the slight thrust misalignment of the center engine.

S-II/S-IVB separation occurred as expected and without producing anysignifiCant attitude deviations.

Satisfactory control of the vehicle was maintained during first and

second S-IVB burns and during parking orbit. During the Command and

ServiceModule (CSM)separationfrom the S-IVB/InstrumentUnit (IU)and

during the T ransposition, Docking and E jection (T D&E ) maneuver, the

control system maintained the vehicle in a fixed inertial attitude to

provide a stable docking platform.

A fter T ranslunar Injection (T LI) attitude control was maintained for the

propellant dumps and chilldown experiment. For AS-505 the APS proDel-

lants were not: depleted by the last ullage burn, and control was main-

tained until the batteries were exhausted.

If.2 CONTROL SYSTEM DESCRIPTION

The control system was essentially the same as that on AS-503. The

flight program was modified to provide for early S-II Center Engine Cut-

off ( CE CO ).

I I - I



1 1 .3 S -IC CO NT R O LS YS TE M E VA LUA TIO N

The A S -50 5 control system performed satisfactorily during S -IC poweredflight. L ess than 1 5 percent of ava ilable engine deflection was usedalthough the actua l flight wind m agnitude was at tim es close to a95-percentile M ay wind.

A ll dynam ics were well with in vehicle capability. In the region of highdynamic pressure, the maximum angles-of-attack were -3.3 degrees inpitch and 2.8 degrees in yaw. T he m axim um average pitch engine deflectionwas -0 .6 degree and was caused by a wind shear, T he m axim um average yawengine deflection was 0 .6 degree and was due to a wind shear. A bsenceof any divergent bending or slosh frequencies in vehicle motion indicatesthat bending and slosh dynamics were adequately stabilized.

Vehicle attitude errors required to trim out the effects of thrust im -balance, thrust misalignment, and control system misalignments were wellwithin predicted envelopes. Vehicle dynamics at S-IC/S-II first plane

separation were well within staging requirements.

11 .3.1 L iftoff Clearances

T he vehicle cleared the mobile launcher structure well with in the avail-

able clearance envelopes. R eduction of the camera data showing liftoffmotion was not performed for the AS-505 flight, but simulations with

flight data show that less than 20 percent of the available clearance

was used. The ground wind was from the southeast with a magnitude of

8.2 m/s (16.0knots)at the 18.3-meter(60-ft)level.

The predicted and measured misalignments, soft release forces, winds,

and the thrust-to-weight ratio are shown in Table ll-l.

II.3.2 S-IC Flight Dynamics

The control parameter maximums for the period of S-IC burn are listed in

Table ll-2. The pitch, yaw, and roll plane time histories during S-IC

boost are shown in Figures ll-l, ll-2, and ll-3. Dynamics in the region

between liftoff and 40 seconds resulted primarily from guidance commands.

During the period from 40 to ll5 seconds, maximum dynamics were caused

by the pitch tilt program, wind magnitude, and wind shears. Significant

dynamics due to wind shears occurred in pitch and yaw between 70 and lO0

seconds. Dynamicsbetweenll5 secondsand S-IC/S-IIseparationwere

caused by high-altitude winds, separated airflow aerodynamics, CEC O, and

tilt arrest. The prominent pitch attitude error at ll9 seconds may becaused by the loss of fin stabilizing action due to separated airflow.The transient at CECO indicates that the center engine cant was -O.l

degree in pitch and -0.15 degree in yaw.

1 1 - 2



T able II -I. A S -50 5 M isa lignm ent and L ifto ff Co ndition s S um m ary

PARAMETER PREFLIGIITPREDICTED LAUNCH

PITCH YAW ROLL PITCH YAW ROLL

Thrust Msalign- +0.34 4_0.34 ±0.34 0.073 -0.038 -0.049

ment, deg*

CenterEngine .... 0.I -0.15 -

Cant, deg

Servo AmpOff- _0.I ±0.I ±0.Iset, deg/eng

Vehicle Stacking ±0.29 _0.29 0.0 0.022 -0.017 0.0

and P ad M isalign-

ment, deg

Attitude Error -0.034 -0.035 -0.003at Holddown

A rm R elea se,

d e g

PeakSoft 316,000 (71,000) 391,000 (88,200)R elease F orce

P er R od,N (Ibf)

Wind 95 Percentile Envelope 8.2 m/s (16.0 knots)a t 1 8 .3 m eters (6 0 ft)

Thrust-to-_eight 1.197 **Ratio

*T hrust misalignment of 0 .34 degree encompasses the center engine cant.A positive polarity was used to determine minimum fin tip/umbilicaltower clearance. A negative polarity was used to determine vehicle/GSE

clearances.**Data not available for update.

At Outboard Engine Cutoff (OECO), the vehicle had attitude errors of-0 .48, 0 .0 5, and -0 .2 degree in p itch , yaw, and roll, respectively.

T hese erro rs are required to trim out the effects of th rust im balance,

o ffset C enter of G rav ity (C G ), thrust vector m isalignm ent, and con tro l

system m isa lignm ents. T he maximum equiva lent th rust m isa lignments were0 .0 7 , -0 .0 38, and -0 .0 49 degree in p itch , yaw, and ro ll, respectively.

T here was no sign ifican t sloshing observed. T he engine response to the

observed slosh frequencies showed that the slosh was well w ith in the

capabilities of the contro l system .

11-3



S_TBEGINYAWMANEUVER _i7MAcH l_TEND YAWMANEUVER 'q/MAX QS _7B EG IN P IT CH A ND R OL L M A N EUVE R S _T S -IC CE CO:_70UTBOARDENGINECANT S_/TILT ARREST'_TENDROLLMANEUVER _S-IC OECO

M E A S URE D

,-, SIMULATED

*L_,J,--

F-- 0 "_

,_ o_ "_ .. ACTUALATTITUDE

_ '-'-F"-"_ __ -40 _ _ d_z _ COMMANDEDTTITUDE- _" ---

(__ _-" LLI "-=----.

o- _ :: -80

ua_ 1

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-1

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0.8

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(.-_L_---_ I

w_ -0.80 20 40 60 80 I00 120 140 160

P_NGE TIM E, SECOND S

Figure 11-1. Pitch Plane Dynamics During S-IC Burn

1 1 - 4



BEGINYAWMANEUVER _ MAXQENDYAWMANEUVER _ S-IC CECOBEGINPITCH AND ROLLMANEUVER_/ TILT ARREST

S_7OUTBOARDNGINECANT _ S-IC OECO

E N D R O L L M A N E UV ER_7 M A C H 1

MEASURED

> o_ .... SIMULATED

w _ - r_') CT aL ATTITUDE

_°"_ ;/_/ _w'-'COMI4ANDEDTTITUDEE 0rl _ --'-

_ 0.8 /_',

kL_ _1_._0

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- 0 . 40 20 40 60 80 100 120 140 160

R AN GE TIME , S EC ON DS

F igure 1 1 -2. Y aw P lane Dynam ics D uring S -IC B urn

1 1 - 5



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Table 11-2. Maximum C ontrol Parameters During S -IC Boost F light

PITCH PLANE YAW PLANE ROLL FLANE

PARAMETERS UNITS RANGE RANGE PJ_NGE

MAGNITUDE TIME MAGNITUDE TIME MAGNITUDE TIME

(SEC) (SEC) (SEC)

Attituderror deg 1.3 94,4 1.2 II.3 -I.2 14.5

Angular Rate deg/s -0,8 78.9 -0,5 12.7 1.4 15.3

Averageimbal deg -0,6 81.6 0.6 85.5 -0.1 79.2

Angle

Angle-of-Attack deg -3,3 84.1 2.8 86.9

Angle-of-Attack deg-N/cm2 10.81 84.1 8.90 86.9

Dyn am ic P re ss ur eProduct

Normal m/s2 0.443 84.2 0.345 85.6

Acceleration

The normal accelerations observed during the S -IC burn portion of flight

are shown in Figure ll-4. The pitch and yaw plane wind velocities and

angles-of-attack are shown in Figure ll-5. The winds are shown both asdetermined from balloon and rocket measurements and as derived from the

vehicle Q-bali.

ll.4 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 deflection

revealed that vehicle structural bending and propellant sloshing had

negligible effect on control system performance. The maximum values of

pitch control parameters occurred in response to IGM Phase l initiation.

The maximum values of yaw control parameters occurred atS-II CECO. The

maximum values of roll control parameters occurred in response to S-IC/

S-II separation disturbances. The response at other times was within

expectations,exceptfor a pitch rateof -0.6 deg/swhich occurredatthe end of the artificial tau guidance mode.

! l - 7



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RANGE TIME, SECONDS

0 20 40 60 80 100 120 140 160

R AN GE T IM E, S EC ON DS

Figure ll-5. Pitch and Y aw Plane Wind Velocity and Free-S tream

A ngl es -o f-A tt ack Du ri ng S -IC B urn



Table 11-3. Maximum Control Parameters DQring S -II Boost Flight

PITCHPLANE YAWPLANE ROLLPLANE

PARAMETERS UNITS RANGE RANGE RANGEMAGNITUDE TIME MAGNITUDE TIME MAGNITUDE TIME

(SEC) (SEC) (SEC)

Attituderror deg -2.1 207.6 -0.5 464.5 -l.l 165.5

AngularRate deg/s 1.2 208.5 -0.25 207.5 l.l 166.5

AverageGimbal deg -o.g 215.I -0.3 464..4 0.2 167.1A n g l e

The maximum control parameter values for the period of S-II burn are

shown in Table ll-3. Between the events of S-IC OECO and initiation

of IGM, these commands were held constant. S ignificant events occur-

ring during that intervalwere S-IC/S-IIseparation,S-II stageJ-2

engine start, second plane separation, and Launch Escape T ower (L ET)

jettison. The attitude control dynamics throughout this interval indi-

cated stable operation, as shown in Figures ll-6, ll-7, and ll-8.

Steady state attitudeswere achievedwithin 20 secondsfrom S-IC/S-II

separation. The maximumcontrolexcursionsfollowingS-IC/S-IIsepara-tion occurred in the roll axis.

At IGM initiationthe FCC receivedTVC commandsto pitch the vehicleup.

During IGM, the vehicle pitched down at a constant commanded rate of

approximately-O.l deg/s. The transientmagnitudesexperiencedat IGMinitiationwere similar to those on AS-504,

A steady stateyaw attitudeerror of approximately-0.04degree occurred

following S-II engine start. At S-II CECO an additional steady state

yaw attitude error of -0.2 degree appeared. Peak transient after CECO

was -0.5 degree and occurred at 465 seconds, This yaw error occurred in

response to the loss of the compliance deflection of the center engine

cutoff. The center engine was not precanted to allow for compliance

deflection. This compliance effect occurred in the yaw plane because of

the locationof the fixed links. Consequently,the outboardengines

were deflectedin yaw after CECO to compensatefor theyaw attitudeerrorand to stabilizethe vehicle. The deflectionsof the outboardengines

in pitch after CECO occurred later and were a result of a pitch up guid-ance command. This command was generated to compensate for the effect

on the trajectory due to loss of center engine thurst.

ll-lO



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'_7S-IC/S-II SEPARATIONCOMMAND _71GM PHASE2 INITIATED,'_7 S -II S E CO NDP L A N E S E P A R A T IO N CO M M AN D S -II L O W E M R SHIF T

_71GMPHASE INITIATED _7S-II OECO'4_/S-II CECO

=_ 1.0ow M'EASURED'ATA_ NOTEIMULATED

w_ _ 0.5 DATA IS ZERO

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I60 200 240 280 320 360 400 440 480 520 560


Figure ll-8. R011 Plane Dynamics During S-II Burn

11-13



Simulated and flight data are compared in Figures ll-6, ll-7, and ll-8.

The major differences were as follows: steady state yaw attitude error

caused by early CECO , which reflects a lower compliance than predicted

for the center engine; initial transients in the roll axis, which could

be attributed to uncertainities in thrust buildup of the J-2 engines;

and steady state attitude errors, which were caused by engine location

misalignments and thrust vector misalignments.

ll.5 S-IVB CONTROL SYSTEM EVALUATION

T he S-IVB T VC 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

experiencedat S-II/S-IVBseparation,guidanceinitiation,EngineMixtureR atio (EMR ) shift, chi freeze, and J-2 engine cutoff. These transients

were expected and were well within the capabilities of the control system.

ll.5.1 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 ll-9, ll-lO, and ll-ll,

respectively. The maximum attitude errors and rates occurred at IGM

initiation. A summary of the maximum values of critical flight control

parameters during first burn is presented in Table ll-4.

The pitch and yaw effective thrust vector misalignments during first burn

were +0.33 and -0.38 degree, respectively. A steady state roll torque

of 14.1N -m (I0.4 Ibf-ft), clockwise 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 N-m (40 Ibf-ft) clockwise.

II.5.2 Control System Evaluation During Parking Orbit

The coast attitude control system provided satisfactory orientation and

stabilizationof the S-IVB/CSMin the parkingorbit. APS engines Ip and

IIIIv responded on an average of one pulse every 40 seconds during _teady-

state operation of the LH2 C ontinuous Vent S ystem (CVS ), which indicates

that the CVS-induced moments were nose up, nose left, and clockwise

(assuming fin position I down and posigrade orientation). APS engine Iprespondedan averageof one pulse every4 secondsduring02/H2 burner

operation, which indicates that the induced moment was nose up. Pitchattitude control during parking orbit is shown in Figure ll-12. The

data of the figure show only the first 2600 seconds of parking orbit

since there were no significant perturbations beyond that point.

I1-14





_7 S -IVB E S CS_/ IGMPHASE3 INITIATED_/ BEGINCHI BARSTEERING_7 B EGIN CHI F R E E ZE

_ S_/S-IVB FIRST ECO

_-__ 0 -'_--,_

.-.J0 C._

_ LLJ <...)

1v-- I Lt.l Q_ "_

o _ _ w 530 550 570 590 610 630 650 670 690 710 730


Figure 11-11. Roll Attitude Control During S-IVB First Burn

T able 1 1 -4. M aximum Control P arameters During S -IVB F irst B urn


PARAMETERSUNITS RANGE RANGE RANGEM A G N IT U D E T IM E M A G N IT U D E T IM E M A G N IT U D E T IM E

(SEC) (SEC) (SEC)

Attitude Error deg +2.0 563.5 -0.8 573.0 0.6 600.0

Angular Rate deg/s -0.9 565.1 -0.2 562.4 -0.6 554.5

AverageGimbal deg 1.4 563.2 -0.6 573.7 ......Angle

11-16



" _BEGIN _NEUVER TO LOCAL HORIZONTAL

_ _LH2CVSONw +4.0

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-2.0 ^

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m 0.0-_- '*

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<_ 700 800 900 I000 II00 1200 1300 140C 2200 2300 2400 2500 2600_ _ RANGETIME, SECONDS

00:15:00 00:20:00 00:40:00

_ NG E T IM E, H OU RS:M IN UT ES:SE CON DS

F igure 1 1 -1 2. P itch A ttitude Control During P arking O rbit

11.5.3 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 Figures ll-13, If-14, and ll-15,

respectively. The significant events are indicated in each figure. Themaximum attitude errors and rates occurred at IGM initiation and EMR shift.

A summary of the maximum values of critical flight control parameters dur-ing second burn is presented in Table ll-5.

The maximum pitch and yaw effective thrust vector misalignments during

second burn were +0.57 and -0.45 degree, respectively. The steady state

roll torque during second burn was 16.8 N -m (12.4 Ibf-ft) clockwise.

The pitch actuator trim position changed distinctly at EMR shift and at

chi bar guidance mode initiation. The trim position change was approxi-

matelyO.l degree in the retractdirectionfollowingEMR shift and O.ldegree in the extend direction at chi bar, No change in yaw actuatortrim position was evident at either of these times.

The pitch trim position change at EMR shift has been observed on previousflights and is attributed to compression in the area of the gimbal due to

the increased thrust. T his compression requires the actuator to shorten

to return the thrust vector' to its original position. The trim position

change at chi bar is attributed to a sudden change in thrust vector mis-

alignment since there was no thrust change at this point.

l!-17



T able 1 1 -5. M aximum Control P arameters During S -IVB S econd B urn


PARAMETERS UNITS RANGE RANGE RANGEMAGNITUDE TIME MAGNITUDE TIME MAGNITUDE TIME

(SEC) (SEC) (SEC)

AttitudeError deg 2.2 9219.5 -I.7 9343.0 -0.9 9244,0

AngularRate deg/s -I.3 9220,2 0.6 9344.5 0,2 9347.0

Averageimbal deg 1.3 9219.0 -l.O 9343.5 ........

Angle

II.5.4 Control System Evaluation after S-IVB Second Burn

The coast attitude control system provided satisfactory orientation andstabilizationof the S-IVB/CSMafter S-IVB secondburn. APS engines

IIIp and IIIIv fired in response to induced moments from the LH2 CVS.

The difference in the polarity of the induced moment in the pitch plane

can be attributed to a normal change in the location of the engine CG.

The LH2 CVS operation was terminated at approximately I0,451 seconds;

therefore, there was minimal operation of APS subsequent to the TD&E

maneuver. The S-IVBwas controlledduringS-IVB/CSMseparation,docking,and ejection. Pitch attitude control after S-IVB second burn is shown

in Figure ll-16.

11-18



_7 S -.IVB S ECO ND ES C

_7 IG M P HA S E 4 IN IT IA T E D_7 E MR S HIF T

B E G IN C H I B A R S TE ER IN GS -]V B S E CO N DE CO

3 r

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o_o ._j _.,_a,--,m,_-&_

I--0 C_

I--- m

9170 9210 9250 9290 9330 9370 9410 9450 9490 9530 9570

RANGETIME, SECONDS

,_SF , _ , _ ,_7

02:33:00 02:35:00 02:37:00 02:39:00

R AN GE TIM E, H OU RS: M IN UTES: SECON DS

F igure 1 1 -1 3. P itch A ttitude Control During S -IVB S econd B urn

1 1 - 1 9





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S ECTIO N 12

SEPARATION

1 2.1 S UM M A R Y

S-IC/S-II first plane separation was satisfactory. The S-IC retro motorsperformed as expected. S -II second plane separation was satisfactory.S-II/S-IVB separation was nominal. The S-II retro motors and S-IVB ullagemotors performed as expected.

Conwnandand S ervice M odule (CS M ) separation from the L aunch Vehicle (L V)occurred as predicted during translunar coast. The Transposition, Docking,and E jection (T D&E ) m aneuver occurred as expected. A ttitude contro l o f theLV was maintained during each separation sequence.

12.2 S-IC/S-II SEPARATIONEVALUATION

1 2.2.1 S -IC R etro M otor P erformance

T he S -IC retro motors performed as expected and provided for a successfulseparation of the S -IC and S -II stages. T he telemetered chamber pressure

data were high as on previous flights. T he data, when biased according

to previous analyses, showed a slightly lower than nominal chamberpressure.

1 2.2.2 S -II U llage M otor P erformance

The S -II u llage motors performed as predicted providing satisfactorypropellant seating for S -II engine start.

12.2.3 S-IC/S-II Stage Separation

S-lC/S-II separation and associated sequencing were accomplished as planned.Dynamic conditions at separation were well with in staging limits. L ongi-tudinal oscilla tions were observed after separation, as they were on

A S -50 4. N o problems were caused by the oscillations.

12.3 S -II S E CO ND PLA N E S E P A R A T IO N E VA L UA T IO N

S -II second plane separation occurred as predicted. T here were noobservable vehicle dynamics caused by second plane separation.

1 2 - I



12.4 S-II/S-IVB SEPARATIONEVALUATION

12.4.1 S -ll R etro M otor P erformance

The S -II retro motors performed satisfactorily and provided a nominalS-II/S-IVB separation.

12.4.2 S-IVB Ullage Motor Performance

The S -IVB ullage motor performance was as expected during staging,maintaining propellant seating for engine start.

12.4.3 S-II/S-IVB Separation Dynamics

S-II/S-IVB separation and associated sequencing were accomplished asplanned. Dynamic conditions at separation were well with in staging limits.The separation conditions were very sim ilar to those observed on previousf l i g h t s .

12.5 S-IVB/IU/LM/CSM SEPARATIONEVALUATION

T he separation of the CS M from the launch vehicle was accom plished asplanned. T here were no large control disturbances noted during theseparation.

1 2.6 L UN AR M O DUL E DO CKIN G A N D E JE CT IO N E VA LUA TIO N

T he attitude of the L V was adequately m ainta ined during the docking of theCS M with the L unar M odule (L M ). T he L M was then successfully ejected from

the L V by the CS M . T here were no sign ificant control disturbances duringthe e jectio n.

1 2 - 2



SECTION 13

ELECTRICAL NETWORKS

1 3.1 S UM M A R Y

The A S -505 launch vehicle electrical systems performed satisfactorilythroughout all phases of flight. O peration of the batteries, powersupplies, inverters, E xploding B ridge W ire (E B W ) firing units, switchselectors and interconnecting cabling was normal.

1 3.2 S -IC S T A GE E L E CT R ICA L S Y S T E M

B oth B attery N o. 1 (O perationa l) and B attery N o. 2 (Instrum entation)

voltages remained with in performance limits of 26 .5 to 32 vdc duringpowered flight. Battery currents were near predicted and below them axim um lim its of 6 4 am peres for B attery N o. 1 and 1 25 am peres forB attery N o. 2. B attery power consum ption was well with in the ratedcapacities of 6 40 and 1 250 am pere-minutes for B atteries N o. 1 and N o. 2,respectively, as shown in T able 1 3-I. E lectrical shorts were experi-enced on bus ID20 (Battery No. 2) from 170 to 173 seconds and 183 to

199 seconds. Current drain from the battery was not excessive and had

no effect on tape recorder playback. Similar shorts have been ex-

perienced after separation on AS-501 and AS-502.

T able 1 3-I. S -IC S tage B attery P ower Consumption

POWER CONSUMPTION*BUS RATED

DESIG- CAPACITY PERCENTOF

BATTERY NATION (AMP-MIN) AMP-MIN CAPACITY, , , , ,

Operationalo,l IDlO 640 27.3 4.3

Instrumentationo.2 ID20 1250 303.3 24.3

*Operational battery power consumption was calculated from power trans-fer until S-IC/S-IIseparation.

Instrumentation battery power consumption was calculated from power

transfer through 7 seconds of tape recorder playback.

1 3 - I



The seven measuring power supplies remained within the 5 ±0.05 vdc limitduring powered flight.

A ll switch selector channels functioned correctly and a ll ou tputs were

issued with in their requ ired tim e lim its in respon se to com m an ds fromthe IU.

T he separa tion and retrom oto r E B W f_ring un its were arm ed and triggered

as program ed. C harg ing tim es and vo ltag es were with in th e requ irem ents

of 1 .5 seconds fo r m axim um allowable charging tim e and 4.2 +0 .4 vo lts

fo r the allowable vo ltage level.

T he com m and destruct E B W firing un its w ere in the requ ired sta te o freadiness if needed.

1 3.3 S -I I S T A G E E L E C T R IC A L S Y S T E M

A ll battery bus vo ltages remained with in specified lim its throughout the

prelaunch and fligh t periods and bus currents remained with in requ ired

and predicted lim its, M a in bus curren t averaged 38 am peres during S - ICboost and varied from 50 to 55 am peres during S -II boost. Instrum enta -

tion bus curren t varied from 52 to 55 am peres during S - IC and S -I I boost.

Recirculation bus current averaged 95 amperes during S-IC boost and ig-

nition bus current averaged 30 amperes during the S-II ignition sequence.

Battery power consumption was well within the rated capacities of thebatteries as shown in Table 13-2.

T a ble 1 2-2. S -II S tage B attery P ower C onsu m ption

BUS RATED POWER CONSUMPTION* TEMPERATURE

DESIG- CAPACITY PERCENTOF

BATTERY NATION (AMP-HR) AMP-HR CAPACITY MAX MIN

Main 2DII 35 7.86 22.5 311.5°K305.9°K

(lOIOF) (gl°F)

Instrumentation2D21 35 12.1 34.6 312.0°K 304.8°K

(102°F) (89°F)

Recirculation 2D51 30 5.59 18.6 304.0°K 300.9°K

No.l (87.5°F)82.0°F)

Recirculation 2D51 30 5.63 18.8 303.1°K299.5°K

No.2 and (86°F)79.5°F)D 6 1

*Power consumption calculated from -50 seconds.

13-2



T he fifteen temperature bridge power supplies and five 5-vdc instrumenta-

tion power supplies all performed within acceptable limits. T he five L H2recirculation inverters which furnish power to the recirculation pumpsoperated properly throughout the J-2 engine chilldown period.

A ll switch selector channels functioned correctly and all outputs wereissued within their required time lim its in response to commands fromthe IU. P erform ance of the E B W circu itry for the separation system wassatisfactory. F iring units charge and discharge responses were with inpredicted time and voltage limits. The commanddestruct EBWfiring unitswere in the required state of readiness if needed.

1 3.4 S -IVB S T A GE E L E CT R ICA L S Y S T E M

T he three 28-vdc and one 56-vdc battery voltages, currents and tempera-tures stayed well within acceptable limits as shown in F igures 1 3-Ithrough 1 3-4. E lectrica l performance was not affected by the anomaly

reported in paragraph 8.7 since A ft B attery N o. 2 responded properly tothe load dem and placed on it. B attery tem peratures rem ained below the

322°K (1 20 °F ) lim its during the powered portions of flight (th is lim itdoes not apply after insertion into orbit). T he highest tem perature of31 6 .5°K (IIO °F ) was reached on A ft B attery N o. 2, Unit I, after S -IVBfirst burn cutoff. B attery power consum ption is shown in T able 1 3-3.

A ll switch selector channels functioned correctly and all outputs were

issued within their required time limits in response to commands fromthe IU.

T he three 5-vdc and nine 20 -vdc excitation modules all performed within

acceptable limits. T he L O X and L H2 chilldown inverters which furnishpower to the L O X and L H2 recirculation pumps performed in a satisfactorymanner and met their load requirements.

T able 1 3-3. S -IVB S tage B attery P ower Consumption

RATED POWERONSUMPTION**CAPACITY PERCENTF

BATTERY (AMP-HRS)* AMP-HRS CAPACITY

ForwardNo. 1 300.0 150.24 50.1

ForwardNo. 2 24.75 28.47 115.0

Aft. No.1 300.0 84.37 28.1

Aft. No.2 75.0 39.72 53.0

*R ated capacities are minimum guaranteed by vendor.**A ctual usage for 29 ,0 0 0 seconds (0 8:0 3:20 ) based on flight data.

S tage design lifetime is nominally 6 hours 30 minutes.

1 3 - 3



(_) |R/4tSFERI3KTER_t)_

$S1_/E_R OFE (_) FWDBATTO.I _TER CV_-ELelITI} OFF

(_) SSB/FHNT__ (_) _O eAtt Ne 2 I_AtERCYCLEF_ ACTUAL

(_) two8Att NO. I .EAr[RCY_-E_It I) _ _ _T[_ CTCLI_[COflt[N_ED) -- ----EXPECtE_ _C_II_

_0 6AtTNO.2 HEAtEA_O-E _C(eet_LE_l.It_] ;

!,li i̧ ¸ I_l _ "" --', -- I_

_' L.a _1_. i

! ' I I _, i I i ,

I I 1 I 1 I I I I 1 I I l l

I / I0¢ aO [ ....... I ....... I ....... I ........ I ...... I ........ I ...... I ...... I ........ I ........ I ....... I ...... I ...... ]: : 0 _:sO:_O 01:_3:_0 01:_:40 0_:30:_0 03:03:_0 03:3_:40 04:10:00 04:43:20 Os:ls:_ Os:SO:O0 0_:23:20 06:_:40 07:_0:00 08:03:20

_._IGEI_, _mu_s:HItlUTE$:S[C_D$

Figure 13-I. S-IVB Stage Forward Battery No. l Voltage and Current

{_)mm_ _A_tvSYSTEMO._ OFt (_)rU_VE _Ar_OV_ROS[IIONFF ..... _ QAX_

(_ m_[NVIRTERaC_OWEROFF 0 FOrWAROATtERY0.aO_FLeTION .... exeeCt_NOMI_

_ANG._[M[,mOO S_CO,_S

-_:1_:40 c0:16:_0 Oo:sO:¢O 01:23:z0 oI:s_:_O _2:_e:Oo 03:03:_0 03:36:_0 _4:10:00 0_:43:_0 Os:I_:40 _5:_0:00 _:_3:20 o_:s_:_ 07:_0:00 _8:03:a0

_i_ r[_. _rs: ,U_TES:SECO_OS

Figure 13-2. S-IVB Stage Forward Battery No. 2 Voltage and Current

13-4



-- ACTUJU.

.... _oOATA

-EXPECTEOY4ZZ(AL(_) T_JSF[RO IrJT£_AL

(_) AF_eATrXO.I _IT Z H[ATERYCLE _ ) _GIN[ STP_TOK _ AFTSArT_0. I UXlTI _At_r CYCL[ AC[_PT_UE_MZrS(_) E_INESTArt

(_) HEATER_O-I_G _ E_INECUTOFF _ A_ CArTNO. I _lt _ _At[R _Y_E0 _HGINEUTOFF

(_) BUyeRaEP_SS _ AFTBATI_. I NEAreR_CLmG _ AF_eA_ _0. 2 _IT 1 HEATE_VOLe_ST_t or PASSlV_tlO_

(_) A_ a_ _0. 10_IT I (_) AFT_ATT 0. 1 _AIERC_¢l[ _ A_T_Alt _0._ _ lt Z _ATERCV_L_ FUELEAOXeEalME_ _ _F_e_Tt_0. 1 _ Z _EXTErC_CLIN_IE R ¢_C LE

I o _ 2 3 ._ 5 6 7 _ 9 Io 2 ? _"

i r I i T

-_:16:40 00:16:40 _:_:_ 01:23:20 01:_:40 _:30:C0 03:03:_ 03:36:40 _:10:_ 04:43:_ 05:16:40 OS:_:_ _ :23:Z0 _;$5;_ 02;30:_ _;03._

Jb_l_TIN£. _URS;MIJ_UT[S:EC_05

Figure 13.-3. S-IVB Stage Aft Battery Ho. 1 Voltage and Current

T_SF[E TO[_T[_ _ _ LOAD,LIMITS_ NOT_LY

LOX& LH CBILL_ _S (_F _ HEATERYCI*[S

(_) _XlLI_V XVO_/4_.I¢_ OIF u ACleWL

(_) LOX& LH_¢_ILL_W_e_ (_ . A_CEelA_LIMITS

_ , __.r--r_,._-.._m_,, _ _._,_,.,,.,L_==_._=__

' I "ll FI II 11I:ll -1°;EF-LL ,_o HI II I I ¢--iL_ II I li II I_.,,t-_lI I I / I I I I°o! !',1_, I1 , !1 ! ! I!FJ° ,I,. ! II [ L[_,._,!i t" _ ,! ,, /, ,! ,! ! /, /

_GE tll_, _ooosEconDs

_ tlp_. _rs: :_Ir_UYES:E¢_

Figure13-4. S-IVB Stage Aft BatteryNo. 2 Voltageand Current

1 3 - 5



Performance of the EBW circuitry for the separation system was satisfac-

tory. 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 needed.

13.5 INSTRUMENT UNIT ELECTRICAL SYSTEM

Voltages on all three batteries showed a gradual rise as the flight pro-

gressed. This voltage increase is expected as battery temperatures in-

crease. A ll battery voltages remained within normal limits. Battery

currents remained normal during launch and coast periods of flight. The

expected peaks in battery currents were observed during launch with av-

erage currents near predicted levels. The 6D40 battery, which had the

highest average current drain, experienced the greatest temperature in-

crease to 340.0°K (152.3°F ). B attery temperature, however, remained

within normal limits. 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 volt power supply maintained an output voltage 55.6 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.

Voting circuits in the Emergency Detection System (E DS) Distributor all

performed nominally. There is no evidence to indicate deviations in

the other distributors or network cabling.

No forced reset commands were issued to the switch selector indicating

that all commands to the switch selector were received properly and no

com ple men t com man ds were ne ces sar y.

Table 13-4. IU Battery Power Consumption

_TED POWERCONSUMPTION* ESTIMATED*

CAPACITY PERCENTOF LIFETIME

BATTERY (AMP-HRS) AMP-HRS CAPACITY (HOURS)

6DlO 350 225.8 64.5 16.4

6D30 350 219.4 62.7 16.9

6D40 350 350.9 I00.3 >I0.6"*

*BasedonI0.6hoursof availablelightdata.

**CCSlossreportedtll.2hours,

1 3 - 6



...... ,Jl,..

........li i....E9

........ ' , lI0

- 27 ::;:; : .... :=:::J::! i ,: :I i ;

I I i i L[ L....

4°11__1 _ : !'!1!111!!!1! i ,_ ,_, ,' ' I I _'II I" II

-30 _ [ , i ] _ , I I ' '

" /i ]_il /l!il5 :, "

} I" [ " I "13033o, i i ; I I If 1 , I ._o32o ,. i, i I ] ' ! ! : -_1o°., , i J _1_1

N 31o, I , i ti{il!l 4 -lOO_:.,_. , i _|!):I}I;_ ]!_iqJlil/iJJll]:_ °

290 i ' l ---,!

2800 '0.5 1 2 4 6 8 lO 12 14 16 ]8 20 22 24 26 2_: 36 38::44 46 48 _"

_OPERATING LIMIT RANGETIME,lO00SECONDS------ PREDICTED

ACTUAL i I I I ! I I _..I n i I00:33:20 02:46:40 05:00:00 07:13:20 12:13:20

Ol:40:00 03:53:20 06:06:40 lO:O0:O0 13:20:00


F igure 13-5. Battery 6DlO Voltage, C urrent

and Temperature



3 1 I i. . ., : H

28 : : ----

° :

27 -"_ i+..... F ::i::

26 ,,,!,_......_T , li I"l !i',i ':' I ! I II

40_,.i I " iI _io_ 351 ' ' ' i ' ' _ 'E I

30i ..... ]

c_ u 201-i I i I t_, -I

oo 15 ......

330 L I [ I I I ] , ,'_T- <' -130

310 I -llO°,, I_ -IooC

I Lii

i . i -90

_ Ii11! "- :-" _,0 '- I' I i 'I -,0 _"='2_°!111 i : -_0-6°-"='.'5 1 2 4 6 I.'141 182022 4 2 3( 444648_ OPER ATING LIMIT

---PREDICTED RANGETIME,lO00SECONDSACTUAL

L _ _ _ _ i _ _ _ _ i00:33:20 02:46:40 05:00:00 07:13:20 12:13:20

Ol:40:00 03:53:30 06:06:40 lO:O0:O0 13:20:00

R AN GE T IME , HO UR S : MI NUT ES : SE CO NDS

Figure 13-6. Battery 6D30 Voltage, Current and

Temperature





S E CT IO N 1 4

R AN GE S AF E TY A ND CO M M A N DS Y S TE M S

1 4.1 S U M M A R Y

Data indica ted tha t the redundan t S ecure R ang e S afety Com m and S ystem s

(S R S C S ) on the S -IC , S -II and S -IV B stages were ready to perform the irfunctions properly on com m and if flig h t con ditions during the laun ch

phase had requ ired veh icle destruct. T he system properly sa fed the

S -IV B S R S C S on com m and transm itted from B erm uda (B D A ). T he perfo rm ance

o f the C om m and and C om m unica tions S ystem (C C S ) in the Instrum ent U n it (IU )

was satis factory, except during the tim e period from 23,6 0 1 seconds(0 6 :33:21 ) when C C S downlink signal streng th dropped sharp ly until 25,0 9 7

seconds (0 6 :58:1 7 ), w hen the an tenna was switched to the om ni m ode.

T he drop in sig na l s tren g th is su spected to be a m alfun ctio n in the

directional antenna system.

1 4.2 S E C UR E R AN GE S AF E T Y C O M M A NDS Y S T E M S

Telemetered data indicated that the command antennas, receiver/decoders,

exploding bridge wire networks, and destruct controllers on each poweredstage functioned properly during flight and were in the required sta te o f

readiness _f fligh t conditions during the launch phase had requ ired veh icle

destruct. Since no arm/cutoff or destruct commands were required, all dataexcept receiver signa l strength remained unchanged during the flight. P ower

to the system w as cu t o ff a t 7 1 5 .3 secon ds, by grou nd com m an d fro m B D A ,

thereby deactiva ting (sa fing) the system . B o th S -IV B stage system s, the

on ly system s in operation a t th is tim e, responded properly to the sa fingcommand.

R adio F requency (R F ) perfo rm ance aspects o f the system are discussed in

paragraph 1 9.5.3.1 .

1 4.3 C O M M AN DA ND CO M M UN IC A TIO N S S Y S TE M

T he perfo rm ance of the com m and section o f the C C S w as sa tisfacto ry. A

tota l o f 51 known com m ands, consisting o f 234 com m and words, were attem ptedfrom the ground sta tions as shown in T able 1 4-I w ith 21 0 words being

accep ted by th e C C S . T ra nsm iss ion o f the 24 w ords no t accep ted w as

attem pted either when the com m and subcarrier was off a t the sta tion , o rou t of lock onboard the veh icle .

14-I



T able 14-I. Command and Communications S ystem Commands History, A S -50 5

RANGETIME TRANSMITTING NUMBEROFSECONDS HRS:MIN:SEC STATION COMMAND WORDS REMARKS

9555.0 02:39:15.0 Redstone AmbientRepress.SystemOff 4 Not Transmitted

and Cryo On

9565.5 02:39:25.5 Redstone AmbientRepress.SystemOff 3 Acceptedand C rye On

15,232.7 04:13:52.7 Goldstone SwitchAntennaLow Galn l Accepted]6,935.0 04:42:15.0 Goldstone EnableTime Base 8 l Accepted16,952.2 04:42:32.2 Geldstone LH2 Tank ContinuousVent Valve 6 Accepted

C lo se O n

16,955.0 04:42:35.0 Goldstone LHR TankContinuousVent Valve 3 AcceptedCl6seOff

]7,095.5 04:44:55.5 Goldstone LH2 TankRepress.ontrolValve 3 AcceptedOpenOff

17,096.8 04:44:56.8 Goldstone Terminate I Accepted

17,097.3 04:44:57.3 Goldstone LH2 TankRepress.ControlValve 3 AcceptedO pe n O ff

17,098.7 04:44:58.7 Gotdstone Ambient Repress.On andCryo Off 3 Accepted17,135.3 04:45:35.3 Goldstone S-IVB Ullage Engine No. 1 On 3 Accepted17,136.7 04:45:36.7 Goldstone S-IVBUllageEngineNo. 2 On 3 Acceptedi7,152.0 04:45:52.0 Goldstone LOXRepress. Control Valve OpenOn 3 Accepted]7,185.3 04:46:25.3 Goldstone Aux. Hydreullc PumpOn 3 Accepted17,299.0 04:48:19.0 Goldstone Passlvatlon Enable 3 Accepted17,299.8 04:48:19.8 Goldstone Engine He Control Valve OpenOn 3 Accepted

17,300.6 04:48:20.6 Goldstone EngineMainstegeValveOpen On 3 Accepted17,301.4 04:48:21.4 Goldstone 8 SecondLead D_ Words 31 Accepted,No CRP's

17,309.3 04:48:29.3 Goldstone Terminate l Accepted17,309.6 04:48:29.6 Goldstone EngineMainstageValveOpen Off 3 Accepted

17,3]0.4 04:48:30.4 Goldstone Terminate l Accepted17,310.7 04:48:30.7 Goldstone PrevalvesCloseOn 3 Accepted17,311.5 04:48:31,5 Goldstone Engine He Control Valve OpenOff 3 Accepted17,324,0 04:48:44.0 Goldstone Engine He Control Valve OpenOn 3 Accepted17,324.7 04:48:44.7 Goldstone DummyWords 42 Accepted, NoCRP's17,335.5 04:48:55.5 Goldstone Engine He Control Valve OpenOff 3 Accepted]7,336.3 04:48:56.3 Goldstone Passtvation Disable 3 Accepted17,354.3 04:49:14.3 Goldstone LOXTank Repress. Control Valve 3 Accepted

O pen O ff

17,355.7 04:49:15.7 Goldstone LH2 Tank Repress.ControlValve 3 AcceptedO pen O n

17,357.2 04:49:17.2 Goldstone PrevaIvesCloseOff 3 Accepted

17,385.1 04:49:45.1 Goldstone AmbientRepress.SystemMode 3 AcceptedSelectorOff and CryoOn

17,407.3 04:50:07.3 Goldstone PassivatlonEnable 3 Accepted17.408.7 04:50:08.7 Goldstone Engine IgnltionPhaseControl 3 Accepted

ValveOpen On17,410.1 04:50:10.1 Goldstone EngineHe ControlValveOpen On g Accepted17,414.4 04:50:14.4 Goldstone S-IVBU11ageEngineNo. I Off 3 Accepted

17,415.8 04:50:15.8 Goldstone S-[VBUllageEngineNo. 2 Off 3 Accepted17,457.7 04:50:57.7 Goldstone Engine Ignition Phase Control Valve 3 Accepted

OpenOff17,459.1 04:50:59.1 Goldstone EngineRe ControlValveOpen Off 3 Accepted

17,494.9 04:51:34.9 Goldstone LH2 ContinuousVentOpen On 3 Accepted17,496.3 04:51:36.3 Goldstone LH2 ContinuousVent ReliefOverride 6 Accepted

OpenOn

17,499.I 04:51:39.1 Goldstone LH2 ContinuousVentOpen Off 3 Accepted17,500.6 04:51:40.6 Goldstone LH2 ContinuousVent ReliefOverride 3 Accepted

OpenOff

19,741.7 05:29:01.7 Goldstone S-IVBU11ageEngineNo. 1 Off 3 Accepted19,743.5 05:29:03.5 Goldstone S-IVBUllageEngineNo. 2 Off 3 Accepted

19,790.6 05:29:50.6 Goldstone Aux.HydrauilcPumpOn 3 Accepted

19,926.6 05:32:06.6 8oldstone Aux.HydraulicPumpOff 3 Accepted24,051.8 06:40:51.8 Goldstone Set AntennaHigh Gain 3* Accepted25,096.7 06:58:16.7 Goldstoae Set AntennaOmni 3* Accepted

35,994.2 09:59:54.2 Goldstone Set AntennaLow Gain 4 Not Accepted

36,033.2 10:00:33.2 Goldstone Set AntennaLow Gain 4 Not Accepted36,426.0 ]0:07:06.0 Goldstone Set AntennaLow Gain 4 Not Transmitted36,502.0 10:08:22.0 Goldstone Set AntennaLow Gain 4 Not Transmitted36,671.0 10:11:11.0 Goldstone Set AntennaLow Gain 4 Not Accepted

*Oneword Is normallyrequiredto switchantennas. These commandswere repeateddue to missed verification

pulsesat the g_oundstationbecauseof noisytelemetry.

1 4 - 2



A com m and attem pted at 9 555 seconds (0 2:39 :1 5) from the ship R edstone wasnot transmitted because two-way lock with the vehicle had not been

established and the subcarrier was off. T he command was successfullytransmitted at 9566 seconds (0 2:39 :26).

Commandsto switch the CCStransmitter to the high-gain directional

antenna at 24,0 52 seconds (0 6 :40 :52) and to om ni antenna at 25,0 9 7seconds (0 6 :58::1 7 ) had to be repeated because the station fa iled to

capture the address verification pulses and the computer reset pulses.T hese pulses were m issed because of no isy telemetry data, which was due

to the low downlink signal problem discussed in paragraph 19 .5.3.2.S ignal strength data indicate these commands were accepted on the firsttransmission.

Commands to switch the CCS transmitter to the low-gain directionalantenna after 35,99 4 seconds (0 9 :59 :54) were not received on board because

of low signal strength. T he low signal strength was caused by the extended

range and low vehicle antenna gain at th is tim e. T he CCS transponder track-

ing threshold is approximately 20 decibels better than the command sub-carrier threshold. Consequently, although two-way lock was being maintainedduring part of this tim e period, 70 -kilohertz subcarrier lock was neverestablished.

14-3/14-4





The maximum angle-of-attack dynamic pressure sensed by a redundant Q-ball

mountedatop the launchescape towerwas 0.76 N/cm2 (l.l psid) from about84 to 87 seconds. This was only 34 percent of the EDS abort limit of

2.2 N/cm2 (3.2psid).

15.2.4 EDS Event Times

All EDS related switch selector events and discrete indications occurred

as expected and are shown in Tables 15-2 and 15-3 respectively.

Table 15-2. EDS Related Event Times

R A N G E

TIME, TIME FROMBASE,FUNCTION STAGE SECONDS SECONDS

StartfTl -- 0.6 --

MultipleEngineCutoffEnable S-IC 14.5 Tl +14.0

LaunchVehicleEnginesEDS IU 30.5 Tl +30.0Cutoff Enable

S-ICTwoEnginesOutAuto- IU 134.2 Tl +133.6Abort Inhibit Enable

S-ICTwoEnginesOutAuto- IU 134.3 Tl +133.8A bo rt Inhi bit

ExcessRate(P,Y,R)Auto- IU 134.6 Tl +134.0Abort Inhibit Enable

ExcessRate(P,Y,R)Auto- IU 134.7 Tl +134.2Abort Inhibit and Switch

Rate Gyros SC Indication "A"

T2 (CenterngineCutoff) -- 135.3 --

Q-BallPowerOff IU 152.3 T2+17.0

T3 (OutboardnginesCutoff) -- 161.7 --

S/C Controlof Saturn Enable IU 708.9 T3 +5.0

S/C Controlof Saturn Disable IU 8629.5* T6 +0.3*

S/C Controlof Saturn Enable IU 9555.8 T7 +5.0

S-IVBEngineEDS Cutoff S-IVB 16,936.0 T8 +0.2No. 2 Disable

*Calculated Value

1 5 - 2



Table 15-3. EDS Associated Discretes

DISCRETE RANGEIME

MEASUREMENT DISCRETEVENT SECONDS

K73-602 O n E DS or Manual C utoff o f L V Engines A rmed 30.0(Switch Selector)

K74-602 O n E DS or Manual C utoff of L V Engines A rmed 31.3

(Timer)

K81-602On EDSS-ICOneEngineOut 135.3

K82-602On EDSS-ICOneEngineOut 135.3

K57-603Off Q-BallOff Indication(+6D21) 152.3

K58-603Off Q-BallOff Indication(+6D41) 152.3

K79-602On EDSS-ICTwoEnginesOut 161.9

K80-602On EDSS-ICTwoEnginesOut 161.9

K88-602Off S-ICStageSeparation 162.5

15-3/15-4



S E CT IO N 1 6

VE HICL E P RE SS UR E A ND A CO US TIC E NVIR O N M EN T

1 6 .1 S U M M A R Y

T he in te rn a l, externa l, and base reg ion pressu re enviro nm ents for the

S -IC , S -II and S -IVB stages were m on ito red by a series of differentia l

and abso lu te pressure gag es. T h ese m easurem ents were used in confirm ing

the veh icle externa l, in terna l, and base reg ion design pressure environ-

ments. The flight data were generally in good agreement with the pre-

dictions and compared well with previous flight data. The pressureenvironm ent was well below design levels.

T h e veh ic le in terna l and externa l acou stic environm ent was m onitored by

a series o f m icrophones positioned to m easure both the rocket engine and

aerodynamically induced fluctuating pressure levels. The measuredacoustic leve ls were genera lly in good agreem ent with the lifto ff and

in fligh t predictions, and w ith data from p revious fligh ts. T he spectra lan a lys is o f th e ten (o ne of w h ich fa iled) a dditio na l S - IC in tertank

acoustic m easurem ents has not been completed. P relim inary estimatesindicate acoustic levels exceeded 1 60 decibels.

1 6 .2 S UR FA CE P RE S SU R E A N D C O M PA RT M E N TVE N TIN G

16.2.1 S-IC Stage

E xternal and internal p ressure environm ents on the S -IC stage were re-

corded by 1 2 m easurem ents located on and inside the engine fa irings, aftskirt, in tertank, and forward skirt. R epresenta tive data from a portion

of these instrum ents are com pared, in F igures 1 6 -I th rough 1 6 -3, w ith

the AS-504 flight data and a band consisting of data from the first threeS aturn V fligh ts.

Differentia l press6 re is the difference between measured pressure and

free stream sta tic p ressure (P in t-P am b). P ressure loading is thedifference between structural internal and external pressures

(P in t-P ext) defin ed such tha t a p ositive load in g is in the burstdirection.

T he A S -50 5 S -IC eng ine fairing compartm ent pressure differentia ls, shown

in F igure 1 6 -I , ag ree very well w ith previous flight data .

16-I



i_

L

S3

0

lI

]_

i

"_

_O 'lS

L

SH

lOZO

i

I

-

_Or

NSr

i



E NG IN E C OM P AR TM E NT

0.8 _ - -I .2

,:C

___ _ -0.8_<

bA U t.Ll .r_ 0.4 VEHSTA 4.33 in _L _ J Z

u_ (170.47n.) _ _ uJu_. - ,t "0.4 _ ",-i_

-- >, ;/"-,,_ :0.4-0.4

.... AS-501,AS-502, AS-504

.... AS-503, FLIGHTBAND --AS-505

IN TE RT AN K CO M PA R TM E NT0 .8

_ -0.8 "_

_-_ 18 z "_

_z""'_" 0.4 VEHTA(730.318.55in.) . .'_/_S<t_'___,.,,," "" -0.4 _:'::_u,_,.

" 0 - -" - __- . "0 _-,

__ FO_:

I:l-v .4

-0.40 20 40 60 80 1O0 120 140

R A NGE T IM E, S E CO NDS

Figure 16-2. S -IC Compartment P ressure Differentials

16-3



1

6

S

O

-0

"_ _N

z_

INV3

S

#

_

,

i

_

__'_8S

_

Z

Z1V



T he S -IC eng ine and in tertank compartm ent pressure differentia ls are

shown in F ig ure 1 6 -2. T he A S -50 5 eng ine co m partm ent p ressure differen-

tia l a g rees w e ll w ith p rev iou s da ta . T he de la y in the peak o f th e A S -50 5and AS-504 intertank pressure differential was caused by the slower tra-

jectories of these flights. However, the trends and magnitudes of the

A S -50 5 data show goo d agreem en t w ith prev ious data .

T he in tertank pressure differentia l showed the characteristic drop as

th e veh ic le passed th roug h M ach I. O n the first th ree S atu rn V fligh ts ,

Mach 1 occurred between 60 and 62 seconds, while on AS-504 Mach 1 occurreda t 6 8 secon ds a nd on A S -50 5 M a ch 1 occurred a t 6 7 secon ds.

T he eng ine and in tertank com partm en t pressure lo adings are shown in F igure1 6 -3. T he in tertank com partm en t pressu re loading agrees well w ith pre-

vious data. The AS-505 engine compartment pressure loading agrees inmagnitude and trend with previous flight data. However, the slower tra-jectories flown on A S -50 5 and A S -50 4 delayed the data peak by approxi-

mately I0 seconds.

1 6 .2.2 S -II S tage

T he pressure environm ent on the S -II stage forward skirt was m easured by

1 4 externa l absolute pressure measurem ents and one in terna l absolute

pressure measurem ent.

A p lo t o f the pressure loading acting across the fo rward sk irt w a ll is

p resented in F igure 1 6 -4. T he A S -50 5 fligh t data and postfligh t p re-dicted va lues are presen ted in the fo rm of m axim um -m inim um data bands

(positive va lues denote burst pressure). T he A S -50 1 th rough A S -50 4

flight data bands are also shown for comparison. Both flight and pre-

dicted pressu re loadings were obta ined by tak ing the difference betweenthe respective externa l pressure va lues an d the assum ed uniform in ter-nal p ressure, which was m easured at vehicle sta tion 6 2.2 m eters (2448.8

in .) an d p erip hera l a ng le o f 1 9 1 degrees. T he A S -50 5 fo rw a rd sk irt p res-

sure loadings were well w ith in the design lim its and agreed with pre-

dicted va lues and previous fligh t data.

1 6 .3 B A S E P R E S S U R E S

1 6 .3.1 S - IC B ase P ressures

S ta tic p ressures on the S -IC base heat sh ie ld were recorded by four m eas-

urem ents, two of which are heat sh ie ld differentia l p ressures. R epre-

sen ta tive A S -50 5 data are com pared with A S -50 4 data and a band of da tafrom the first th ree S atu rn V fligh ts.

S -IC base pressure differen tia l is show n in F igure 1 6 -5 as a function o f

altitude. In general, the agreement is good between AS-505 base pressuredata and previous fligh t data .

16-5



MEASUREMENT VEHICLE STATION m (in.)

"DI08-219 63.88 (2515

COMPLETE DATA FAILURE D109-2i9 63.75 (2510

D110-219 63.50 (2500* * QUESTIONABLE DATA D111-219 62.99 (2480

D]12-219 62.74 (2470

POSTFLIGHT PREDICTION D113-219* 61.54(2423

AS-505 FLIGHT DATA EXTE_At D114-219 63.25 (2490

I D115-219 63.25(2490ATA FROM PREVIOUS FLIGHTS D116-219 _:_ 63.25 (2490

D117-219 ** 63.25 (2490

MACH I D118-219 61.98(2440AX Q D119-219 63.75 (2510

S_7 S-IC/S-II FIRST PLANE D120-219_ 63.25 (2400D121-219.* 63.25 (2490

SEPARATIONINTERNAL-- D163-219 62.20 (2449

DESIGN'LIMIT (STATIC PRESSURE_ ' 'VARIES FROM 4.34 TO 4.69 N/cm_ Posw I

(6.3 TO 6.8 psi(1)DEPENDING POSl

ON VEHICLE STATION _= 0.___¢ :90.

POS III I VEH STA 63.88 ml; -

F3.76 73.50m // 63.25m

62.99m/ \_ 62.74m

61.54 m 61.98 m --6 .._

4j 4JX X

v

Or3 I # ¢Z}¢_C_

uJ :== l

-

g7 V7 _0 25 50 75 lO0 125 150 175

RANGE TIME, SECONDS

Figure 16-4. S-II F orward Skirt Pressure L oading

16-6





AS-501,AS-502, .-----AS-504A S-503 FLIGHT BAND . AS-505

A L T IT UDE , n mi

0 5 I0 15 20 25 300.4 i

'_ • -0.4

D0046-106 -0.2

<_ _ -_._ _ VEHSTA2.81m _

o_ _ _.,_ , "" ._,_I.,_ 110.63 in.) _o oo

__ -0.2_

- 0 . 2

0 I0 20 30 40 50 60

A LT ITUDE , km

F igure 1 6 -6 . S - IC B ase H eat S h ie ld P ressu_ L oading

A S -50 5 fligh t da ta band. T hese effects can be attribu ted to w iden ing o f

the da ta band caused by an unpredicted local increase of static pressure

a fter second p lane separa tion . T he gradua l dec_ in hea t sh ie ld a ft face

pressure from second plane separa tio n to S -II O utboard E ngine C utoff(OECO), noted on flights AS-503 and AS-504,.was not as pronounced duringthe A S -50 5 flight.

The predicted pressures after S-II CECOwere based on four- and five-engine 1/25 S-II stage scale model test results. It is shown that the

predicted minimum pressure during this time interval does not follow the

fligh t da ta trend. T h e fligh t da ta sh ow a pressure drop a fter C E C O a t

all m easured locations, while the pressure measured during model testingincreased or decreased depending on location . T hese m odel da ta are no t

strictly app licable, because the effect on the pressure distribution asa result of the center engine shutdown was not simulated.

1 6 .4 A CO US T IC E N V IR O N M EN T

1 6 .4.1 E xtern a l A coustics

T he A S -50 5 externa l fluctuating pressures were m ea sured at s ix veh icle

sta tions loca ted on the S -IC aft sk irt, F in D base, S -IC in tertank, S -I I

fo rw ard and aft sk irts , and S -IV B fo rw ard and aft sk irts. F igure 1 6 -9

16-8



POSTFLIGHT

.... PRFDICTION

A S . - 5 0 5

FLIGHT

DATA _/S-II ESC

X_/S-II SECOND PLANE SEPARATION

I PRI-VIOUS _TS-II CECOLIGHTDATA _EMR SHIFT

_7 S-I I OECO

VEHSTA' '

m 41.4m

I /---I]66 •(85,n.)-

F-__ " J/_- . . D102-206

i,.,-- "_Er IW,_F__LDO99-2O6 o.100.06 L,___- _,_. R;2.16(8sin.)_

-_--_: 90deg _ : 0 degx-__187-206 O.08

" ---- VEHSTA46.0m (1810n.) --" " -_AZIMUTHdeg

. 0.04 J [RADIUS3.30 m (130 in.) -0.06

0.04" MAXIMUM uJ

0.02 "_

•=:,_ _ ,_>MINIMUM 0.02

0 _ _ ==_-_ 0

-0.02

-0.02_7_ _ _

150 200 250 300 350 400 450 500 550 600 650

RANGE TIME, SECONDS

Figure 16-7. Thrust Cone and Base Heat Shield

Forward Face Pressures

16-9







S-IVB FORWARD SKIRT

1401 AS-SO3 !-505"h_ _ S-II FORWARD SKIRT4OIIAS-503 _.LAS-5°4"_ I I lIllll---,_--_--I IoAsPL148.4b ' . ...._-E_3°" ' _,,,,, _IIIIIII I _ ,-_'-'-_IJOASP'!_3.,O#bS.-SpsZt_jLI II III1![_'%_ "1[ ]! !llil "TITllK I B0025-426 _-130

12011 I i i Ill', _l'ii'r'i',_.' VEH STA81.62m _&_E II Jl I i JlIHI I _"_[_'__ I IIIltBO_

_

,_ollI ', ilJJi _ITl I I r'I'h,,3'_ f I liil _ IIo,,_,.,..;.......,_.; ,,,, s- i -1ir.r]_ I ,,',,.i.,

_ loll I IIIIlll] IIIIIII I IIIII1_, I)!tlll,,.,,', I1 I i tt)ll IIIIII ] I lllllrlx\l II ,,-,,__..,, II I lltlllU IIIIIIJ I))1111 --:_',dl[)llli.. ,ooll 1 I i llti IIIIII Illlll \&l III ! =_" 10011I I i llllll IIIIIII I IIIIII r,?,.xlIIHI

_"''" ijAsslosliilt illlll I IIIIIIII N'klllll _',.,.,,.,.,..IAs.slosIIIIIII lilllll I IIIIII; I'_!]11!!11901JOASPL151.4b IIIIII I I l iltlll t "i,'lll II _ 90IloAsPLlS4.'ldb I IIIIII I I I IIIIJ I r'l li l)l l

S-IVBFTSKIRT S-IIFTSKIRT

_ '4011AS'S03 I !!lJ!! I ._ 14011AS-SO3 ' I l|lA_-sost-,Iit)If, ,

' ....... I_ _. = I IOASPL = 148.5 db ;-Sub_ "= >"_"30|] I : I 'IIH II._ i OASP,is3._,b ,AS-'_ I 1,11111[ ,'1

° I[ "" "_ IAs'sl°'iillt' llll)Ji =_ . l_O { , I I I II"I ,"

1 I i i1 [1 1 1 1 J ,I ' I I _ :- , 'V ,, , , , , . I _ Il l t = __z'I_-. 120 i , ill! I I I IIll_ ._''' ''2776t_/...n./ .a--.

, il h IJ:JJJ I J_T_I ii I I)"h_ _ 120

' [I _"7 I IOASPL = 153.7 db _"_AS-503 )'='_,_1664.47 !n...).

_ . ())IIIII l)l)Ir_',J,l)llU ,,,r,oiiiii r ltttliHI I , II1,11t

IIIIUI I!;','s_S !11111 llilll ix.l,ltllll _.. lOO i I l llillI IIIlll I I]l))/ll I I"_!i)= .... _g As-sos

150.9 db ...... I _\ 90' OASPL = 154.9 db lllllJ I I I JJlllJ ! I i!!i

1 10 100 lODD 10,OO l 10 100 1DO0 10,000

F RE QUEN C Y, HERTZ

F RE QUEN C Y, HERTZ

F igure 16-10. Vehicle E xternal Sound Pressure Spectral

Densities at L iftoff ( Sheet l of 2)



S-llAFTSKIRT S-ICAFTSKIRT

140AS-503 _'_=_ _ 15011AS-503 III Illlll <.o_._

" ' ' ' _ TBO002-I

.a_ 120 AS-504 ' i EH STA 42.28 m .... _ .<== II As-5o4 ]liJ_]Y/TN\_,'(12o.o8n.)J

_-_ OASPL = 147.4 db - , 1664.47 in.) OASPL : 154.2 db , !:__11o__ ,oolf i i I II,,, _ff:A , I'_:.L_ ,_x AS-

_ toos-505 _-_ IIAs-soS IIilllllllil,:i;_! ,I:- 90OASPL=I47.8_b :_ IO0[]OASPL15_.0biIfIII I I !: :

S -I C I NT ER T AN K

_= I IOASPL : 150.4 db #( NI ."_ N 150rTAs.50I T" FIN D' >-- l .,' , IOASPL=161.1 db--T- -

_. [ As_5oH-_I, [LLHi'-- II_ I_. BO0-18 ::_ O04-11

_m 120 i AS'5103 _F_--_ _,VEH STA19.05m I__ IIAS-504 I i LrI!,'"L_, " ii,_U_75o.oo..) $ _ 13o_ IIOASPL = 156.8db I I ri_-_u_: _

_ ! iiiili immlli_,x,,[tlI,l _ _

_ 1°°1 i _ ll°_'_ 90,OASPL=I_5.1b IIIIIII I )'_lIIIII _" 100_........... --1 10 lO0 1000 I0,000 1 l0 1O0 I000 10,000

FREQUENCY,ERTZ FREQUENCYHERTZ

F igur e 16-10. Vehicle E xte rn al S oun d P re ss ur e S pectr al

Densities at Liftoff ( Sheet 2 of 2)



S-IVB FORWARD SKIRT $-II FORWARD SKIRT16o ': I I" I I } f 1 1 _ _8c _ I _

,_ _ ;AS-505_I i L _-_AS-5021 J _ ! i i (

_. \_V'_,J, ',.

' i _ VEHTA1.62, ' __ I _ F i I ) ) ? ) ) .

I

"P_ _ 160 S-lVB AFT SKIRT _ I_ S-II AFT SKIRT

_, i ! I m I I I I _. ! I I

_o ,..-._---_As ° 5 _ i _s=_.-t14o __ _I_-%_" . =_. ASSS 8-200_ "_I _ _ _ 1 - ! _ I ! VEH STA 42 28 m

-_ _: _ =_ -_ I 1 I ' -

x I VEH STA 70.52 m _ × I _,"K t -_ I_AS-S04 i i _i :

_'_ i _ _ !'_-liil ii-%'_.2776.37 in.)

•. i i I , , , l _ Izo

lOQ 20 40 60 80 IOO 12O 140 >_ 20 40 60 80 I00 120 140

RANGE TIME, SECONDS RANGE TIME, SECONDS

Figure 16-11. Vehicle External Overall Fluctuating

Pressure Level (S heet l of 2)



S-II AFT SKIRT _ S-IC AFT SKIRT[

L_8o° _=160 I I AS-505-_ _-a B 002-115_

_-_ _ j 37-200 _ .160 VEH STA3.05m_ i I i IAS-53--_ _., VEH STA 42.28 m _ _ { I i I(120.08 in.)r_"- I _ _ _ _ _ (1664.57 in.) _ _ " AS-BnS L- Ac_5O3_ I. I I !

,._, , _ ._. I I I _ "'_._ /1 i_._._----_._u_ × I AS-5O4 _ _ ! i ! AS-504 "_--.--

_,',; 120 I I I _ _ i I '

S-IC INTERTANK _ FIN D

_ 160 ' 1 =.. 16( _ i 1

- ,,I i : _.) _ T_O_. ,,s-_ VE,-,T,,.O_,,

"" I -- t i I li / '_ I i ,i ; i_ I_ 120 _ "" 120_

o_"_ _0 40 60 80 100 120 140 _ _ 20 40 60 80 100 120 140

RANGEIME,ECONDS RANGEIME,ECONDS

Figure 16-11. Vehicle E xternal O verall F luctuatingPressure Level (Sheet 2 of 2)



N 150 S-II )RWARDSKIRT S-ICINTERTANK

_.14cOAFRL156,6b J _::130OAFR'-38b_____/___

_ T = 55.5 sec AS-503 E[] -'_-_i {!IglTII869__ I =_" T-- 84 sec ' :--'.!;. 1 ' , ZIBO003_II __:: 13C AS-504 , -- _ m _ 120 AS-504 :_ ".-" p._ ;T,y-_ • '

o OAFRL = 154.2 db -J NJ \ _" _ 110 ....

l; ,li p lJl "_AS,SPIF]'FEL_ i[IINI __ , , '

5' J'J'"__ " __s-5'o !I,--AS,-SPE,I!ill \ Illllll _.._1_o (_.:o.81 11 I IIIIII]l _ It}lrll _" XOOiM__ii;I}il} i_%i' '_ OAFPL: 155.0 db-- _ I___

lOOT=6o_ Jill f []fltlll _.\!]illll _oiT=85_ _ ',140 S,--,IIAFT SKIRT 5-IC AFTSKIRT

,AS-SO3M_=I.8 1 {I llllll .-LI ( [IT--_'--%_I[ _: 130 AS-503 _ [_"T'_'r _I Isi\i i =o_ 130,OAFPL = 152.2 db i1505 _ _ M.=1.78 _A_: '

T : 79.5 I I k_. --_ t, _oo3z-aoo =_-:_-" 1 'il :H l I i. ._L--_ iNSTA 42.28m =,_.

_ 120_s-504 ' 1664.5zn.) _= 110 AS-SO¢ _ _(I_0.08 io.),,oo_.45 { iF _ ' '", I '-'" ..:_.55

b _' OAFPL : 143.7db I k i"_'_ _\ f I I I [ _ _ OAFPL: 138.5 db i ; JR:_ 110 T :80s_ ! I_/ _-_AS-<. "_\ s!! _ 100 T :82.0s_

5!!!!t!! [i', i l '_liiill _.. 90AS-505 ___ .. 100 I I M. = 1.8 I ii illi,::1.5 Ill , I I_ll!ill .... ,: 'O_ _ _ DAFPL: 147,2db _ iOAFPL= 143.3db-_JF-_-H_

="_9oT=_o_ I t i I I ! t1!iill _0T: 85_oo ' , ,,

N 1401 S-I AFT SKIRT 150 FIN_s-5o_ I IIIIIll I I IIIIII_ _ AS-_O_ I I,( ::

r--__ <_.=1.3 l I II!lit , , , ,, ,,c...i I I Iii111 I l' _=_-'- 14®:1.oz <,,

_,1_o OA_PL:145.8<t_III!111 I Itlli1! /_ _ 140OAFPL:150.gdbT : 71.5 sec I _ T : 62 sec i !

I ! iiH ! I lltllt I t1 1 1 1 1 ' - " A S -6 0 4_ 120AS-504 ,,,,,,i 'i 0 2 . m " AS-50I_' '!l i IL_];_ _:: 130 f VEN STA 2.06m

AS-503_ _ (1664.57in.) _-_

) )TIIIH &_-°_ RIO OAFPL= 142.2db ] _-- 120 OA_PL = 148.4 db ', i

%.Ill,i!, _.....

I I''' _7. _ T : 70 sec i _\ __--AS-505 , l ii85secIll!l t:::I:R"HQI1"_ASTSO4ihi _c'_J, II!ll I. iiiiiii I li: \ _.',t].._,,,._lt_ll,,i...........:< ' !" II111_s-s0511ltlii Ii1111 _ As-50_ I IAS ' ". ' "i i i i _ I i i i , , i i i _ , I

_.. lOOA_-sos _ ,.:o.9_ , i_F:_,_-,' , ......,,,_ ,.:1.8 ! I111111IIIIIIII !1%1!1111_ II0 OAFPL=148.gdb ) II)_'C'=''"": i:I._ OAFPL = 143.0db _

_o T:85,o_ I IIIllll I IIIIIlil I IIIIIII 1ooT:64,_.. I I llll_l"..._, , ,,, , -,.--- .,10 I00 I000 lO,O00 l I0 lO0 lO00 I0,000

FREQUENCYHERTZ FREQUENCY,ERTZ

FiguY_ 16-12. Vehicle External Fluctuating Pressure Spectral

Densities at Maximum Inflight A erodynamic N oise



on the phasingof the loads (pressures). An estimateof the phasingand

pressure amplitudes can be established when the spectral analysis of all

the data is completed.

16.4.2 Internal Acoustics

16.4.2.1 S-IC Stage. Internal acoustics were measured at two locations

on the S-IC stage. One measurement was located in the intertank section,and the other' in the thrust structure above the heat shield. The acoustic

data at these locations are shown in Figures 16-13 and 16-14. Data from

both measurements agree with previous flight data, except AS-505 datawere somewhat: lower between 20 and 60 seconds.

16.4.2.2 S-II Stage. The two internal microphones, used on the S-IIstage, are located on the forward skirt and thrust cone. Figure 16-15

presents the internal overall acoustic levels versus range time for

AS-505. The forward compartment internal acoustics show agreement with

previous flight data during liftoff. The lower level for the aft compart-

ment internal acoustics is due to the measurement commutation occurring

after the liftoff acoustic maximum.

AS-505 internal and external acoustics are shown in Table 16-1 and com-

pared with data from previous flights. The microphone located away from

MAXIMUMSPL _ OVERALLSPL.IMIT

MEASUREMENT PREVIOUSFLIGHTDATA AS-505 LEGEND

B005-106 144.8 at 0 143.7 at 0 169.0 x_--_x_PREVIOUSLIGHTseconds seconds _%_ DATA ENVELOPE

AS-505 FLIGHTDATA

[:::='SPLn db referenced to 2 x 10-5 N/m2

160

.140

o=1 2o , ._......._...---.."< :o--r""i'-

ll I "_'110 0

0 20 40 60 80 1gO 1ZO 140 160

RANGE TIME, SECONDS

Figure 16-13. S-IC Heat Shield Panels Internal Acoustic Environment

16-17



MAXIMUMSPL [_ OVERALLSPL

LIMITMEASUREMENT PREVIOUSFLIGHTDATA AS-505 LEGEND

PREVIOUS FLIGHT001-I18 144.7 at O 139.5 at O 157 7_'_

seconds seconds D AT A E NVE LO PE

AS-505 FLIGHT

DATA

[_SPL in db referenced to 2 x I0-S N/m2

°o13 ',t'

115 1

O 20 40 60 80 100 120 140 160

RA NGE T I_ E, S EC ON DS

Figure 16-14. S-IC Intertank Internal Acoustic Environment

the tower (270 degrees azimuth) shows the higher level, as expected. Thedifferential between the forward external and internal acoustic levels is

approximately 8 decibels at liftoff. The differentials for Mach l and

Max Q conditions are 18 decibels or higher because the greater high fre-

quency contents are more attenuated across the vehicle skin. The differ-ential between the aft external and internal acoustic levels is not

realistic because the internal measurement was commutated after the lift-

off acoustic maximum.

16.4.2.3 S-IVB Sta_e. The S-IVB acoustic environment was measured at

fourpositions,internaland externalon the forwardskirt, and internal

and external on the aft skirt. Both external measurements provided valid

data only during portions of the flight due to an instrumentation mal-f u n c t i o n .

Time histories of the composite levels, 50 to 3000 hertz, for these loca-

tions are presented in Figure 16-16. The AS -505 structural transmissi-bility for the sound pressure at liftoff is indicated by the difference

(shaded band) between the external and the internal measurements. T he

maximum external levels and minimum internal levels measured during the

A S-503 flight are also shown, indicating that the AS-505 levels werenominal.

16-18



7 MACH ] '_ "MAXQ _ S-II ESC

160 DESIGNEVEL40db-J FORWARDCOMPARTMENT

-_ 1 50c.} Z

_ 140

-_ 13o _" /7/ \_ .. \', # \,&.u _ %

ov 120 .t _ \

AS-505-- ---- P R E VIO US F L IGHT S

160 DESIGNEVEL61dbAFT COMPAR'TMENT

i._oa

"at 150_ z

_ 14o

° I-J 130 " _b _x _& |

"" " lJl,., I _,

o_ 120 " I

,,,, ,VV, I, Zz, , I , ,-20 0 20 40 60 80 lO0 120 140 160 180 200 220 240 260


Figure 16-15. S-II Internal Acoustic History

T able 1 6 -I. S ound P ressure L evel Comparison of A S -50 5W ith P revious S aturn V F light Data

MA XIMUM O VE P_ LL DB

FORWARDOMPARTMENT AFTCOMPARTMENT

EVENT EXTERNAL INTERNAL EXTERNAL INTERNAL

(B0]6-219) (B017-219) (B037-200B038-200) (BO3g-206)

AS-505 AS-501/ AS-505 AS-501/502/ AS-505 AS-501/ AS-505 AS-501/502503/504 503/504 503/504 503/504

Liftoff 154.1 154.0 145.7 ]42.0 154.9 153.7 128.7 137.5

Transonic ]53.5 156.5 133.9 133.0 143.5 147.8 129.0

MaxQ ]54.0 151,2 137.0 ]38.0 147.2 152.2 129.0

Max ]39.5 134.0 150.3 161.8

S t a t i c

F i r i n g

16-19



160 l '

FORW ARDSK IRT _1 MEASUREMENTA I L E D

120 _'_

- !._1

DATADROPOUTS_J

-J100 I ___J _ ASo503 AS-505 EXTERNAL-- -- INTERNAL=u_ 16O

_ AFTSKIRT i_ MEASUREMENTFAILED

I O 0

0 20 40 60 80 lO0 120 140 160


Figure 16-16. S-IVB A coustic Levels During S-IC B urn

16-20



S E CT IO N 1 7

VE HICL E T HE RM A L E NVIR ON M EN T

1 7 .1 S U M M A R V

T he A S -50 5 S -_C base region therm al environm ent was sim ilar to that ex-

perienced on earlier fligh ts with the maximum temperatures genera llyh igher as a resu lt o f h igher am bien t tem pera tu res a t lifto ff. H ea t

sh ield temperatures and structural temperatures forward of the heat

shield were generally with in the bands of p rev ious da ta and well below

design a llowan ces. T he fo rward surface and bondlin e m easurem ents didnot indicate loss o f M -31 insu la tion on A S -50 5.

S -IC fuel tan k and in tertank skin tem peratures exceeded the predictedm axim um during the early portion of fligh t. T h is condition w as a resu lt

of the h ig her am bient tem pera tures and wind veloc ity a t lifto ff.

B ase therm al environm ents on the A S -50 5 S -I I stage were sim ila r to those

measured on previous flights and were well below design limits. Heat

sh ield aft surface tem peratures increased between S -II Center E ng ineC uto ff (C E C O ) and E ngine M ixtu re R a tio (E M R ) stepdown , bu t were w ellbelow the design predictions.

T he aeroheating ra tes on the A S -50 5 S -II stage in tersta ge, body structure

and fa iring s were sim ila r to those on prev ious fligh ts.

T he A S -50 5 S -lV B stage aeroh eating env ironm ent was co m parable to th at o fA S -50 1 , A S -50 3, and A S -50 4 and coo ler than that o f A S -50 2.

1 7 .2 S -IC B AS E H E A T IN G A N D S T A G E S E P A R A T IO N E N V IR O NM EN T

1 7 .2 .1 S - IC B ase H eating

T herm a l env ironm ents in the base reg ion o f the S -IC stage were recorded

by 29 m easurem ents, which were located on the heat sh ie ld and F - I engines.

T his instrum enta tion included 6 radia tion calorimeters, 1 6 to ta l calori-m eters, and 7 gas tem pera tu re probes. R epresen ta tive da ta from these in -

strum ents are com pared with the A S -50 2 th ro ugh A S -50 4 flig h t data band.S ee F igures 1 7 -I and 1 7 -2. Data are show n versus a ltitude to m in im ize

tra jectory diffe rences. A S -50 1 fligh t data , wh ich sh owed less severity

than subsequent fligh t data beca use of flow deflector effects, are notshown.

1 7 - I



o

r-[

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V

9

NV

H

N

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17-2



A S -50 5 S -IC base therm al environments have sim ilar trends and magnitudes

as those m easured during the previous fligh ts, as shown in F igures 1 7 -I

and 1 7 -2. In genera l, A S -50 5 radiation hea ting ra tes were slightly

h igher than A S -50 4. M axim um va lues o f radia tion and tota l heating rateoccurred a t a ltitudes betw een 1 5 and 22 k ilom eters (8.1 and 1 1 .9 n m i).

T he m axim um tota l hea ting rate m easured in the A S -50 5 base region was39.5 watt/cm 2 (134.8 Btu/ft2-s), recorded on the inboard surface of engineN o. 3 (C 1 23-I0 3). C E C O on A S -50 5 produced a sp ike in the environm en tswith a magnitude and duration similar to previous flight data at CECO.

AS-505 base gas temperatures show good comparison with AS-502 through

AS-504 flight data. However, AS-505 gas temperature data do not show

the decrease between 4 and 9 kilometers (2 and 5 n mi) that previous

flight data indicated.

The heat shield temperature data are compared to previous flight data in

Figures 17-3 and 17-4. Measurement locations for the S-IC base heat

shield are shown in Figure 17-5. The temperatures were generally higher

than on previous flights largely because of a higher ambient temperature

at liftoff. The forward surface and bondline measurements did not indi-

cate M-31 insulation loss, and temperatures were well below design levels.

Measured temperatures showed reasonable agreement with.a flight reconstruc-

tion (not shown in the figures) based on flight radiation data, gas tem-

perature data, and design heat transfer coefficients.

Engine temperature data were normal. The thermal response of the turbineexhaust manifold under the insulation on the inboard side of engine No. l

at vehicle station -l.l meters (-44 in.) is shown in'Figure 17-6. The

measurement trace is similar to previous flight data. Temperatures under

the insulation on the gimbal actuator and on the fuel discharge line were

well below design limits while gas temperatures inside the engine cocoons

remained within the band of previous flight data.

17.2.2 S-IC/S-II Separation Environment

Forward skirt compartment gas temperatures, shown in Figure 17-7, were

similar to those encountered during separation on previous flights. Two

spikes in the gas temperature were noted. The first spike was due to the

S-If ullage motor flow field and the second spike was due to the five J-2

engine plumes. Peak temperatures, due to the J-2 engine plumes, may

have reached slightly higher peaks than those shown, at approximately

4.0 seconds after separation since data at this point exceeded the

upper limit of the transducers, requiring extrapolation between 3.7

and 4.4 seconds after separation.

17.3 S-II BASE REGION ENVIRONMENT

The S-II base heat shield and thrust cone flight environment was, in

general, in good agreement with previous flight data and postflight pre-

dictions. Base heat shield measured heating rates, gas temperatures,

17-3





FIN C

POS Ill

POS IV

C034-I

FIND. FINB

C038-II5

_,C033-I06

/POS II

-(

POS I

C029-106FIN A

LOCATION CODE

, HONEYCOMBs_. _ - FORWARDSURFACE o

FLIGHT'a'/" F"_. "" "._,,iT l' _2.59 cm LOSSOFM-31 ASNOTED •

' _(I,..02 in.) BY TV CAMERAON

II PREVIOUSFLIGHTS

M-31 .,>_ . .i _I M31/HONEYCOMBNTERFACE "

3-_._I.0.76 cm_(O.3,0 in. )'J"-_:_"_ _ I 0 25 cm

THERMOCOUPLE LOCATION"=J (().l in.) FROMAFT SURFACE 0

Figure 17-5. S-IC Base Heat Shield Measurement Locations

17-5



lOOO

) R E DI C TE D _ X

f- -4 --_ .................. -------... 1200

900 / _ / W

/ / .....

800 / _-- _-_ _'_ _ 1000

I// / PREVIOUSFLIGHT

DATAA,Di _;_%_LIG"TA

700 // / 8006oo /

I//"' ODOr,O400 / /

/ I" 200/

300 _ /

-20 0 20 40 60 80 lO0 120 140 160 180 200

R _GE T IME , S EC ON DS

Figure 17-6. S-IC Temperature Under Insulation,

Inboard Side Engine No. l

AS-505 FLIGHT DATA [ ._-EXTRAPOLATED1400 • C164-120

• C165-120 _" _ 2000PERFO_ANCELIMIT5340°F ._ j- ,...

. . . . . R A ,E L IM IT_ _'_-'OFTRANSDUCER '1800. .% ,.200

_ ',_ 1_oo

_ 1400

1000 UPPERIMITF \-_ ,_"

PREVIOUSFLIGHT DATA "_ J_ ]200 _

800 • \ \ -]ooo

i "_ _" _, 800/'-'_ • "ND

600 "_ / 600

"_ / j// _ ____'_'_ "\ _ ; 400

¥ o2O O

1 2 3 4 5 6 7 8

FII4EAFTER SEPADATIOI_,SECONuS

I I I I I

163 165 167 169 171


Figure 17-7. S-IC Forward Skirt Compartment Ambient Air Temperature

DuringS-IC/S-IIStage Separation

1 7 - 6



thrust cone heating rates and aft face tem pera tures are presented in

F igures 1 7 -8 th rough 1 7 -1 1 , a long with prev ious fligh t da ta and post-

flight p redictions. T he predicted effects o f C E C O o n the heat shield

heating rates were determined from four- and five-engine 1/25 scale S-IIstage m odel test resu lts; o ther predictions were accom p lished by the sam e

analytica l methods described in previous fligh t evaluation reports.

------POSTFLIGHIPREDICTION _7 S-11 ESC

_#7 S-II SECOND PLANE SEPARATION

--AS-5O5 FLIGHT DATA _/ S-If CECO

_7 EMR SHIF T

][ PREVIOUSFLIGHTDATA _7 S-IfOECO

C858-206, R = 1.33 m

( 52 .5 i n. )

.f_. .C719-206C724-.206 .-"-_J/_C717-206

R = 1.91 m/_< _R = 1.35 m (53 in.)

THESECALORIMFTERS (75in.)_

ARE LOCATED FLUSH WITH C718-206_-_-__ _C687-206, R = 1.96 m (77 in.)

THEAFT FACEOF THE R = 2.31m_ _C720-206

6 BASE HEAT SHIELD (91 in.)1 7"-, _ R = 2.11 m (83 in.)

(_: 90de_ _ =Odeg --5

o_ C722-206E 5 R = l.93 mu

(76n.) ?

-_ -MAXIUM -_

uC 4W

,.<, =

2

150 200 250 300 350 400 450 500 550 600


Figure 17-8. S-II Heat Shield Base Region Heating R ates

17-7



TEMPERAUTE,

°K

HEATING

ATES

,

att/

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THESESURFACE C715-206 C710-206

TEMPERATURE R = 2.54 m (I00 in.____/R = 1.35 m (53 in.)MEASUREMENTS AZ = 180 deg_ AZ = 228 deg

ARE.LOCATED C711-206 ___" C683-206

FLUSH WITH R = 1.8 nl(71 in.)/__-'_--_.,_,,_,, ,, 1.90 (75 in.)

THEoFHEAFTHEATFACE AZ = 82 deg_._)-}--- AZ== 278 mdeg

SHIELD _I/I , C713-206

|!_?_!_i'_i_i_|ESIGN PREDICTION _ = 9fl°_'TF R == 1350deg.32(52 in.)PAST FLIGHT DATA LC714_206AZ

I\\\\\\N AS-505 FLIGHT DATA R = 2.21 m (87 in.)AZ = 0 deg

llO0_ 1400

v 800 _f

°

=_ 700 " 800

400 S_7S-IIESC

S-II SECOND PLANE SEPARATION

S-IfCECO 200

EMR SHIF T

300 _/ S-II OECO

I I I o00

100 _7_7 '_

150 200 250 300 350 400 450 500 550 600

R ANGE TIME, SECONDS

Figure 17-11. S-II Heat Shield Aft Face Temperatures

17-9



As expected, an increase in the base region environment after CEC0

was noted, except for the heat shield maximum heating rates. After

CEC0 these heating rates continued at approximately the same level of

magnitude until engine mixture ratio shift and did not experience the

predicted increase as shown in Figure 17-8. At the lower heating levels,

an increase in total heating rate of approximately 0.6 watt/cm2 (0.5

Btu/ft2-s) was observed, consistent with the predicted values. Thrustcone heating rates, shown in Figure'17-9; base gas temperatures, shown in

Figure 17-10, and heat shield aft face temperatures shown in Figure 17-11,exhibited slight increases after CEC0.

1 7 .4 VE HICL E A E R O HE A T IN G T HE R M AL E N V IR O NM EN T

17.4.1 S-IC Stage Aeroheating Environment

A erodynamic heating environments were measured with therm ocouplesattached internally to the structura l skin on the S -IC forward sk irt and

intertank. Generally, the aerodynamic heating environments were higherthan fo r the A S -50 4 fligh t but were below design lim its.

Measured skin temperatures and derived heating rates for the S-IC inter-tank are shown in F igure 1 7 -1 2. P ostflight s im u la tions of sk in tem pera -tures and heating ra tes are a lso presented. T hese sim ulations are based

on analytica lly determ ined heat transfer coefficien ts and recovery tem -

pera tures u ntil flow sep ara tion reaches th e in tertank. During the perio d

of flow separation , a radiation heating environment, determ ined from pre-vious flight data (AS-502 and AS-503), is used in the simulation. Good

corre la tion was obtained between the fligh t data and the sim ula tions.

T he S -IC forward sk irt sk in tem p era tures and derived heating ra tes are

presented in F igure 1 7 -1 3. T he A S -50 5 S -IC forward sk irt skin tem pera -

tures and derived heating rates were higher than recorded on AS-504. TheS -IC fo rward sk irt was u n insu la ted on both A S -50 4 and A S -50 5.

Flow separation on the AS-505 flight, according to ALOTS data, occurred

at approximately ll6 seconds. The forward point of flow separation ver-

sus flight time is plotted in Figure 17-14. The effects of CE C0 on the

separated flow region during AS-505 flight were the same as observed on

AS-503 and AS-504. It should be noted that at higher altitudes, the

measured location of the forward point of flow separation is questionable

because of loss of resolution in the flight optical data.

L0X tank skin temperatures were well below the predicted maximum through-

out flight, as shown in Figure 17-15. There was a noticeable measurement

response when the L0X level passed corresponding thermocouples.

F uel ta nk sk in tem peratures exceeded the predicted m axim um during the

early portion o f fligh t, as show n in F igure 1 7 -1 6 . T he h igher in itia l

tem peratures are attribu ted to h igher am bient tem perature and wind velo-

city a t lifto ff. T hese tem peratures were with in design lim its.

17-10



340 , ' I_i 15o

36°L!'32o100 _

/" _ _ 3oo _._._.L__ I _l /" L _

_ 50__ 280

:_iilto 80 12o 16o i I i ib r _ ' " i

RANGEIME,SECONDS 0 I , I0 4O 80 120 160

FIN C FIN B FLIGHTDATA RANGE TIME, SECONDS

__ _AS-SOI, AS-502

_'_'=_ANDAS-503 FIN C_ / FIN B

FIN _ FLIGHTATAFIN A --'--AS-504

_ _--STA 724 AS-SOS _=:=_STA_ 1491 --.--AS-504....AS-SO5 POST- FIN Dk_. _ FIN A AS-505

C63-I18_ PLIGHT _FSIMULATION C64-120J

2 2

I i

0 -:m--.-_ , . _ 0 _ 0 -- _ _ 0_ _...... >_-. _ -,_

= N-1 -1

0 40 80 120 160 0 40 80 120 160

R AN GE T IME , S EC ON DSR AN GE T IME , S EC ON DS

Figure 17-12. S-IC Intertank Figure17-13. S-IC Fo_ard SkirtAerodynamiceating Aerodynamiceating



N OT E: F OR WA RD L OC AT IO N

OFFLOWSEPA_TION 2000

50 MEASUREDFROMAS-505

ALOTS 70 _ FILM

S-ICCECO

40 1500c

Z0 Z

C _ I000_

o _ )= 20

lr n 50 0

olOO I10 120 130 140 150 160

_NGE TIME, SECONDS

Figure 17-14. Forward Location of Separated Flow

Intertank skin temperatures exceeded the predicted maximum during the

early portion of flight, as shown in Figure 17-17. This condition was

due to the higher ambient temperature and wind velocity at liftoff.

The forward skirt skin _temperatures were slightly higher than on AS-504

and were considerably higher than the first three flights, which had in-

sulation on the forward skirt of the S-IC stage, as shown in Figure 17-18.

However, temperatures reached a maximum of 319°K (ll5°F ) at the end of

the flight which was well below the predicted maximum.

17.4.2 S-II Stage Aeroheating Environment

S-II stage aeroheating data were in good agreement with previous flight:

data and postflight predictions and were within design limits. Measured

heating rates for the aft interstage, ullage motor fairing, LH2 aft and

forward feedline fairings, forward skirt and systems tunnel forward fair-

ing are shown in Figures 17-19 through 17-24.

Flight data from measurement C863-200, located on the forward fairing of

ullage motor No. 6, are lower than data from C861-200 at a similar loca-tion on the forward fairing of ullage motor No. 4 and the predicted heat

rates, as shown in Figure 17-20. Ullage motor No. 6 is located in the

vicinity of the LOX vent valve. Cold gases from the LOX vent valve in the

form of either a cooled boundary layer or allowable leakage from the vent

valves, or a combination of both, could be flowing over the calorimeter

17-12



I 20O

350 I f

PREDICTED MAXIMUM _, / I00

\i

300 / AS-505 FLIGHT DATA BAND/,, _ C173-I19 C177-119

/ /f C175-I19 C179-119

o_ / // C176-I19 0

. 250 /"

_'_-- I /1 r "-- ........... _zoo " "'" _ ,_7_,7_, -10o_

I- f/'/ DATA BAND i.,t. _ "I" _ _ 1-

1oo_ _ _ ..... .................. -_oo

,o I0 20 40 60 80 100 120 140 160


Figure 17-15. S-IC LOX Tank Skin Temperature

400 I

240

380 /

PREDICTED MAXIMUM _ /

\ _,/ 200

360

o /" LAJ

/ / 16oI--

/ /

__ AS-S05 FLIGHT DATA BAND _ _-C080-117 COBl-ll7 / / -120

320 / // / i" "'--

300 -- ---"-"

_'_ _ _./" _ 80

............... -_.,_....,_.__._.--...... _ .... _. PREVIOUS FLIGHT

280 DATAAND

I I I ,o0 20 40 60 80 IO0 120 140 160


Fiigure17-16. S-IC Fuel Tank Skin Temperature

17-13





PREDICTION _FLIGHT DATA 1PREVIOUS- -- -- FLI GHTDATA3 .0

MACH1 VEHSTA43.5 m(1715.75 in.)

2.5- S_7S-IC/S-IISEPARATON 0

VEH STA 43.2 m (1701.Iin.) _IIllll

COMMAND AZ = 10 deg / l lJll ll lll ll --2.02.0 C909-200 J lllllllI]lll

VE H S T A 42.0 m (1 654 in .) III I II I II II IZ = 10 degqE u_

] '.5

3= :m4-_

--I0_

1.0 ,,,

Z

_ 05 .

az MAXIMUM-k J _ r

M I N I M U M J

-0.5

-i.o 27 _ _70 25 50 75 I00 25 150 175

R AN GE T IM E, S ECO N DS

F igure 1 7 -1 9 . S -II A ft In terstage A eroheating E nvironm ent

disk and thus indica ting a reduced heating ra te. T he data from C 86 1 -20 0

are h igher than the predictions during the period o f I0 5 to 1 25 seconds;

th is difference is presently under investigation.

L ower than expected hea ting ra tes were recorded by ca lorim eter C846 -21 8

o n the L H2 aft feedline fa ir in g . T h is con ditio n is possibly a resu lt o fcoo l g ases from w ith in the fa iring bein g driven ou t o f th e a ft jo gg le due

to rap idly decreasing pressure outside the fa iring.

A t range tim es grea ter than 1 35 seconds, flight da ta from m easurem ents

C 80 1 -20 1 , C9 0 5-20 0 , and C9 0 9 -20 0 , shown in F igure 1 7 -1 9 ; C 846 -21 8 and

C 847 -21 8, shown in F igure 1 7 -21 ; and m easurem ent C 81 1 -21 6 ,shown in F igure

17-15



POSIFLIGHT I PREVIOUS--PREDICTION --FLIGHT DATA FLIGHTDATA

3"01 tILLAGEMOI:ORFAIRING

TOTALHEATINGRATES 1

VEHSTA42.01m (1654in.)

2.5 ' 1 1 ___.nC861-200 AZ=139 deg _

2.0 C863-200AZ=230 deg @

E AA

MACH _'_

I.oMAX,_ ev

_" _7 S-IC/S-IISEPARATIONCOMMAND I_- _(-_Z _-_

I-'- _

,.<,0.5 _

- 0 . 5

-_.o '_' _ '_'25 50 75 lO0 125 150 175

R AN GE TIME , S EC ONDS

Figure 17-20. S-II Aft Interstage Aeroheating Environment,

Ullage Motor Fairing

17-24 are higher than predicted. At the higher altitude, F-l engine ex-

haust gases are drawn forward along the vehicle surface into the low pres-

sure region created by the separation of the boundary layer. R adiation

from this exhaust gas could cause the higher heating rates indicated bythe calorimeters.

Additional measurements in the form of S-II stage structural, fairing,

and insulation surface temperatures were made during the AS-505 flight.

Data from these measurements (not shown in the figures) agree well with

previous flight data and postflight predictions and were within designlimits.

17-16



POSTFLGHT I PREVIOUS__PREDICTION _FLIGHT DATA FLIGHTDATA3 ,0

S_7MACH

_7 MAXQ

2,5 S-IC/S-II SEPARATION COMMAND

I I -2.oC846-218

_(1782.68 in.)

E AZ=135 deg ,_

° @c847-218 _

._ 1.5" VEHSTA44.35m --

_ (1746n_) _P. AZ=98deg m

" I J

0.5 ;_-'_ _:

-0 .

-I.0 _Z _7 _p'

0 25 50 75 100 125 150 175

RANGE TIME, SECONDS

Figure 17-21. S-II A ft Interstage Aeroheating Environment,

LH2 Feedline Aft Fairing

17.4.3 S-IVB Stage Aeroheating Environment

The mission profile of the AS-505 flight produced nominal thermal environ-

ments for the S-IVB stage components and structure. The thermal severity

of the A S-505 boost trajectory was comparable to that of AS -504, AS -503,

and A S -501, and cooler than that of AS -502 and the thermal design trajec-

tory. There was no instrumentation from which structural temperatures

could be obtained; however, since the thermal severity of AS-505 was less

than that for the design trajectory, the S-IVB stage structural tempera-

tures should be within the design limits for the boost phase.

1 7 - 1 7



x_7MACH1 _?MACH 1

IVMAX IVMAXIV S-IC/S-IISEPARATIONCOMMAND IV S-IC/S-TISEPARATIONCOMMAND

PO5TFLIGHT I PREVIOUS POSTFLIGHT ]"-- PREDICTION --FLIGHT DATA FLIGHTDATA ------PREDICTION FLIGHT DATA PREVIOUSFLIGHTS

VEH6TA 63,4 m (2517 in.)

AZ = 225deg2.5 _ 2., --

E820-219

(_ liiiiiii!il-2.0 AzVEHTA 63.4m (2517n.)=5 deg -2.0.0 "_ 2.C (_ --

LHFEEDLINE iiIZI!_!III C803-219

PA_RINGOTAL i__ii!i_iii VEHSTA62.3 (2473n.)

///_ i _i_i__ ' AZ : 45 deg

1.5 HEATINGRATES .(__ _ _ 1.5

(_)C848-218 _(_ _ _ _

. VEN STA 49.46_ (1947in.) ._ _ _ _:_l f0

AZ:_5*g .7.-'_, _ - -1.o .

. ®DG52-,,o ;EHSTA 48.5 m (1911in.) _ k_ _03 _ AZ = 45 deg _ _ ,"" _ _: 0.5 ,-" _ _ A

--.,_..=_ _ .o - _- -_ -_ -o

-0. -0

-i.o _ IV _ -i.o IV IV0 25 50 75 100 126 150 175 25 50 75 I00 125 150 175

RANGETIME, SECONDS RANGE TIMEo SECONDS

Figure 17-22. S-II Body Aeroheating Environment, Figure 17-23. S-II Body Aeroheating Environment,

LH2 FeedlineForwardFairing ForwardSkirt



POSTFLGHT PREVIOUSAS-503)

-----PR |'DIC T IO N - F LIGHT DA T A -- F LIGHT D A T A

3.0 l , ' FLOW_/ MACH 1 I

'_/ MAXQ ITS-IC/S-II SEPARATIONCOMMAND

2.5

2.0 C811-216 --

VEHSTA 63.01 m (2480.6 in.)AZ = 30 deg _,

E N

u 1.5 I. _ ...

_ .m .IJ

, .o.0 :_

_ i I . _

,,, 0.5 \ ,,<,.._ -r

0

-0.5

-l.O _7 _ _'25 50 75 iO0 125 150 175

RANGETIME, SECONDS

Figure 17-24. S-II Body Aeroheating Environment,

Systems Tunnel Forward Fairing

17-19/17-20



S E CT IO N 1 8

E NVIR ON M EN TA L CO N TR OL S Y ST EM

1 8.1 S UM M A R Y

T he S -IC can ister condition ing system and the aft environmenta l condition-

ing system performed satisfactorily during the A S -50 5 countdown.

T he S -II therm al contro l and com partment condition ing system main tained

temperatures within the design lim its throughout the prelaunch operations.

Available data show that the Instrument Unit (IU) Environmental ControlS ystem (E C S ) perform ed satisfactorily. T he IU env ironm enta l co ndition ing

purge duct exh ibited a pressure loss and flow increase during prelaun ch

operations but IU performance was unaffected.

1 8.2 S -IC E N VIR O NM EN T A L CO NT R O L

T he am bient tem peratures o f the I0 can isters in the S -IC fo rward sk irt

com partm ent m ust be m ain ta ined a t 29 9 .8 ± II.I°K (80 ± 20 °F ); however, the

can is te rs can opera te a t 324.8 to 27 7 .6 °K (1 25 to 40 °F ) fo r no m ore than

a I0 m inu te period. N o can ister condition ing is requ ired a fter S -ICforward umbilical disconnect.

T he am bient temperatures with in the can isters remained with in the requ ired

lim its during "the countdown. C aniste r N o. 1 recorded the lowest tem p era -tu re, 289 °K (6 0 .5°F ), during pre launch. T he lowest canister tem perature

m easu red in fligh t (F igu re 1 8-I) w as 25 9 °K (6 .5°F ) in can is te r N o . 2.

During J-2 eng ine ch illdown prior to lau nch, th e th erm al environ m ent is

at the m ost critica l po in t. W ith in th is period the am bien t tem pera tu re in

the fo rw ard sk irt com partm ent dropped, as show n in F igure 1 8-2. T he low -

est temperature, 185°K (-126.7°F) was recorded at instrument C207-120which is located under a J -.2 eng ine nozzle and received the m axim um effecto f the co ld helium . A ll o ther am bien t tem pera tu res w ere above the 20 5.4°K

(-9 0 °F ) design m in im um . T he band of am bien t tem pera tu res during flight

(F igure 1 8-2) exceeded the predicted m axim um but th is did no t cause a problem .

T he design requ irem ent for the a ft com partm ent is that the prelaunch

tem pera tu re be m ain ta ined a t 29 9 .7 ± 8.3°K (80 ± I5°F ). A ft com partm ent

tem pera tu res are sho w n in F igure 1 8-3. P rio r to L O X loa ding th e a ft

com partm en t tem peratu re w as a m axim um of 31 1 °K (IO 0 .4°F ). T h is is 3°K (5.4°F )

above the m axim um perform ance lim it bu t did no t cause a problem .

18-I



Jo

'3_nlV_3dH31

_o

'3_I_3

dH31



28O III Ill]PREDICTED MAXIMUM "7 AS-505 FLIGHT DATA BAND- c_6_-_2o czo_-_zo: ZB

: ' 1 6 5 - I 2 0 C208-120

_ 220

200 -lOO

.i"_ _ _..'_ //' ,1

,iao-_ 11" -'_.--_._;._I'I PREDICTEDMINIMUM--_ ""--_.._-'_/J"_

,iooI I I I c_o,-[_o-,.....O 20 40 60 80 'iO 'i0 'i0 'i0


F igure 18-2 . S -IC F orward C ompartmen t A mbien t T emperature

1 8 - 3



370. 400

35 0

200

_o_ ---4----t---4--14o _ _ 1 I :

• = T330 AS-50S FLIGHT DATA BAND . i I

C108-I1S C20D-115 _ 1SO%

•_00

....................... 300-2"

Z70

-25,140 -25,120 -25_I00 -25,080 -25,060 -2S,DAD -25,020

RANGE TIME, SECONDS -180 -160 -140 -120 -100 -80 -60 -40 -ZO

RAHG_ TIME, SECONDS

-6:59:00 -6:58:20 -6:57:40 -6:5:00I._ RANGE TIME, HOURS:MINUTES:SECONDS aO

POSIllUMBILICAL

_8 _2D

FI_ ¢ FINB

06 9

' 1 8 0

l.S2m (60 In.) _ 140_WDOF '

POS IV _" _S-SOS FLIGHT DATA BAND '

POS II 3_,0 $107-115 C203-115

I'74\ .100

, . \ /

/ / C_04-II_ TT_RIES 300_HSTA lZKIO

..... \ " 61 m / .....PREDICTED MINIML

--,,CZ,_ I '_ 20 40 60 80 100 120 140 160

POSL RANGE TI/_, SECONDS

F igu re 18-3 . S -IC A ft C om pa rt me nt T em pe ra tu re R an ge





--RTG PURGE DUCTING-D68-603

•_ y\ ',

-.. : PROBABLEOCATION.. .... l OF LEAK

"'-. _..,. ', ', _ ........

"'''-... ii i II , ...+.-"

............. ]_ ......... t_l ................

Figure 18-4. RTG Purge Ducting Modification

18.4.1 Thermal Conditioning System

Sublimator performance and coolant temperature during ascent are presentedin F igure 1 8-5. Im m edia te ly after lifto ff, the M odula ting F low Co ntro l

Valve (M F C V ) began driv ing toward the fu ll hea tsink position wh ich was

achieved at approxim ately 20 seconds. T he water va lve opened at 1 81 seconds

allow ing water to flow to the sublim ator. F ull cooling from the sublim ato r

was not evidenced until approxim ate ly 530 seconds. A t th is tim e, the

coolan t tem perature at the contro l po in t began to decrease rap idly. T he

low coo ling rate during the first 30 0 seconds after the water valve opened

is typica l o f a slow-sta rting sublim ator. A t the first therm al switch

sam pling , th e coo lant tem p erature wa s still abo ve the actuation po in t and

the water va lve rem ained open. T h e second therm al sw itch sam pling occurredat 7 83 seconds and the water valve closed.

T h e IU coo la nt flowrate was sligh tly below the m in im um specifica tion lim its

as shown in T able 1 8-I. T he out-o f-specifica tion flo wrate w ill be eva luated

in regard to applied pum p voltage. N o degrada tion of system perform anceoccurred.

18-6



..................................OPEN -2_ _ ....... L _._d..........,_o_sw _c, -_o288 _ ......SETTEMP287 _'--. ! :

286 _ I285 WATERALVEPEN _- I

FIRSTTHERMAL SWITCH SAMPLING _ ----z--_

284--_7 WATER VALVE CLOSED i 1 5083

u 2.5 SUBLIFt_TORATER 1 I .3

2.0 _ INLET PRESSURED43-601/ I ! "

.1.s _I_._Ly .2_.0 I ' =

° i I.5 i

i //_ 1 -30,000

Io I I ..URLIMATOREAT -- I

REJECTIONATE 20,000

i 1o,oooo _f-- , I o0 100 200 300 400 500 600 700 800 900 1000 1100 1200 1300 1400 1500

_ NGE T IME, S ECO NDS

, _7 , _ , _F , ,0 0:05:00 ():lO:O0 0:15:00 0:20:00 0:2_:00

R A NG E T IM E, HOUR S :MI NUT ES : SE CO ND S

Figure 18-5. IU Sublimator Performance During Ascent

Tab_,e 18-I. TCS Coolant Flowrates and Pressures

MINIMUM MAXIMUM

PARAMETER REQUIREMENT OBSERVED OBSERVED

IU CoolantFlmvrate 6.06 x lO-4 5.87x lO-4 5.99x lO-4

F9-602m3/s (gpm) (9.6Minimum) (9.3) (9.5)

S-IVBCoolant 49.2±2.52x IO-5 4.80x IO-4 4.92x lO-4

(7.8±0.4) (7.6) (7.8)lowrate

FI0-602 m3/s (gpm)

PumpInlet Pressure 10.82 to 11.72 11.17 11.24

D24-601N/cm2(psia) (15.7 to 17.0) (16.2) (16.3)

PumpOutlet 28.89 to 33.23 29.32 30.70Pressure (41.9to48.2) (42.5) (44.5)

D17-601 N/cm2(psia)

1 8 - 7



The TCS GN 2 sphere pressure decay, which is indicative of the GN 2 usage

rate, was approximately as expected for the nominal case as shown in

Figure 18-6. The rapid pressure drop during the first 780 seconds, though

not predicted, is not considered an abnormal condition. The same type of

drop was present during the AS-504 boost phase, and is regarded as a

cooling effect of compartment outgassing during boost.

T he Flight C ontrol Computer (F CC) contains an internal coolant flow pass-

age which is normally connected in parallel with the TCS flow loop. Due

to a potential failure of coolant tube connecting flares inside the FCC

cover, the IU-505 FCC was disconnected from the TCS flow loop. Thermal

vacuum test performed prior to launch showed that with no internal cooling

the upper allowable temperature limit of the FCC would not be exceeded

under worst case hot conditions. The predicted worst case temperature

and available flight data are presented in Figure 18-7. The internal

temperature remained well below the allowable and predicted worst case

temperature limits.

Component temperatures appear to be within the expected ranges, but

insufficient data preclude any conclusive comments at this time

(F igure 18-8). L imited real-time information and second-burn data indicate

all component environmental parameters were satisfactory.

18.4.2 Gas Bearing Supply System

The gas bearing subsystem performed nominally through the time period for

which data are available. The GBS GN2 sphere pressure decay is nominal as

can be seen in Figure 18-9. Figure 18-10 shows the platform pressure

differential and internal ambient pressure. The platform internal

pressure(D12-603)decreasedas expectedto 8.63 N/cmZ (12.5psia)at

4000 secondsthen increasedto 9.80 N/cm2 (14.2psia) at 24,000seconds,

however the gas bearing differential pressure ( Dll-603) exhibited the

expected tendency to increase during the initial portion of the flight,

which has been seen in most previous flights. Data after mission com-

pletion show the differential pressure steady and below the maximum allow-able value.

18-8



22- 3200

21- 300020-

19- 2800

18- 2600

17- 240016-

% 15- 2200

14- 2000_

x 13-

12- 1800

_ll- 1600m

LulO"_: 1400-

12O0CIC

(:"7" lO00

6 -

5- 800

4- i600

3" 400

2 -

l' 200

00 2 4 6 8 lO 12 14 16 18 20 22 24 26 28 30 32 34 36 38 40 42 44 46 48 50

R ANGE TIME , lO00 S EC ON DS

i i I i I I i I i I ! , I I I i

l:O0:O0 3:00:00 5:00:00 7;00:00 9:00:00 11:00:00 13:00:00


Figure 18-6. Thermal Conditioning System GN2 Sphere Pressure (D25-601)

353 ; _ _ , l I ' I

, '. i i I i i _ : " : , , J-170i ! J i MAXIML_4ALLOWABLE TEMPEPJ_TURE= 373°K (212°F); ' : ' ' , _160

343 : : i I ; I ... I _ I ; ' . i I E 1 , IREDICTEDMAXIMUMWORST CASEHOT CONDITION _ : .

;:;} ! : I ! : : Fo333_ I i . 140

_ : 1 _ i : : : 130 _3

323 ' IJ "_- ] i I I , i ..._ ' '. 1-120 _

_ _ _ 0

313 -lO0

• _ i J: i /! , ,

• ..,..1L_ _/" ! i : ; 90

! ACTUALEMPERATURE 80

i '93 , i f i " ' ' " ' I i i -7o, i i i i i I ,,i i ; , , # , , -

2 4 6 8 lO 12 14 16 18 20 22 24 26 28 30 32 34 36 38 40 42 44 46 48 50

R AN GE TIME , IDO 0 S EC ON DS

i 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 ll:O0:O0 13:00:00

R A NG E T IM E, HO UR S :M IN UT E S: SE C ON DS

Figure 18-7. Flight Control Computer Temperatures (C69-602)

18-9



5

.r_

EI

:"

o

_

odr1.2

.

7

_

_

o

,

.



22i 3200

• 3000

20 "2800

18 2600

2400

16'

2200

14 2000

× 1800_-

1600

,_0. I1400_

_1200a_8 T

_: 1OOOL.

68O O

4 600

4O O2

20 0

0 0

0 4 8 12 16 20 24 28 32 36 40 44 48

R AN GE T IME , 1000 S EC ON DS

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


Figure 18-9. Gas Bearing GN2 Sphere Pressure

I I I I I I I I 17

GA S B EA RIN G PR ES SUR E DIF FE RE NT IA L

II.5 - DII-603

I 11 SPECIFICATION I i/ / /

1 4

13 I I I I I I I 19

- PLATFORMNTERNALMBIENT 18

12- PRESSURE(D12-603)

17

II 16'_

z 1B/ I

: _ _ F[IGH_PERATi.GRAN_E 121a

7 1 1 1 1 1 1 1 1 10

0 4 8 12 16 20 24 28 32 36 40 44 48

R AN GE T IME, 1000 S EC ON DS

i t I I I I I I I I I I I

2:DO:O0 4:00:00 6:00:00 8:00:00 I0:00:00 12:00:00

R AN GE T IM E, H OUR S :M IN UT ES : SE CO NDS

Figure 18-10. Inertial Platform GN2 Pressures

18-11/18-12



S ECTIO N 19

DA TA S Y ST EM S

1 9 .1 S UM M A R Y

A ll elements of the data system performed satisfactorily except for aproblem with the Com mand and Communications S ystem (CCS ) downlink duringtranslunar coast.

M easurement performance was excellent, as evidenced by 9 9.2 percentreliability. T his is the highest reliability attained on any S aturn Vf l i g h t .

Telemetry performance was nominal, with the exception of a minor

calibration deviation. The onboard tape recorder performance was

satisfactory. Very High F requency ( VHF ) telemetry R adio F requency (R F)

propagation was generally good, though the usual problems due to flame

effects and staging were experienced. VHF data were received to 15,780

seconds (04:23:00). C ommand systems R F performance for both the S ecure

Range Safety CommandSystems (SRSCS)and CCSwas nominal except for theCCS downlink problem noted. G oldstone (G DS ) and G uaym as (G Y M ) reported

receiv ing CCS signal to 40 ,1 9 1 seconds (1 1 :0 9 :51 ). G ood tracking data

were received from the C-B and radar, with B erm uda (B DA ) indicating fina lL oss of S igna l (L O S) a t 35,346 seconds (0 9 :49 :0 6 ).

T he 73 ground engineering cameras provided good data during the launch.

1 9.2 VE HICL E M EA SUR EM E NTE VA LUA TIO N

T he A S -50 5 launch vehicle had 2286 measurements scheduled for flight.F ifteen measurements were waived prior to the start of automatic countdown

sequence leaving 2271 measurements active for flight. O f the waivedmeasurements, 2 provided valid data during flight.

T able 1 9 -I presents a sum m a_ of measurem ent perform ance for the tota lvehicle and for each stage. M easurement performance was exceptionallygood, as evidenced by 9 9 .2 percent reliability, which is the highestattained on any S aturn V flight.

T ables 1 9 -2, 1 9 -3, and 1 9 -4 tabulate by stage the waived measurements,totally and partia lly failed measurements, and questionable measurements.N one of the listed failures had any significant impact on postflightevaluation.

1 9 - I



T able 1 9 -I. A S -50 5 F ligh t M easurem ent S um m ary

S-IVB

MEASUREMENTSS-IC S-II STAGE INSTRUMENTTOTAL

CATEGORY STAGE STAGE PHASEI* PHASEII _ UNIT VEHICLEii i

Scheduled 669 1018 378 378 221 2286

Waived 3 9 2 2 1 15

Failures 5 6 4 8 0 19

Partial Failures 22 13 5 5 0 4Q

Questionable 0 4 0 0 0 4

Reliability, 99.2 99.4 98.9 97.9 lO0.O 99.2

Percent

*Notes: I. S-IVB Phase I period of performance is from liftoff to parkingo rbit i ns er ti on .

2. S-IVB Phase II period of performanceis from liftoffuntil end

of S-IVBstagerequired flight period of performance as

specified in the Detailed Flight Test Plan.

Table 19-2. AS-505 Flight Measurements Waived Prior to Launch

MEASUREMENT

_UM_ER MEASUREMENT TITLE NATURE OF FAILURE REMARKS

S -IC S TA GE

C343-I15 LOX Prevalve, Engine 5 Data negative KSC waiver I-8-505-4

Dllg-I04 GimbaI System FilterManifold Transducer failure MICH-50S-4

D128-I15 LOX Suction Line, Engine 2 i(olsydata prior to MICH-50S-3, Date satisfactorylaunch, during flight.

S -If S TA GE

C758-217 LOX Tank Liquid Tem_)erature Transduceropen

C850-218 E4 LH2 Feedli_eHeatRate Transduceropen

{3030-201 LH2 RecircPump Disch Pressure Transducerfailure Installation

D030-202 LH2 Recirc Pump Disch Pressu_ Transducer failure installation

i D030-203 LR2 Recirc Pump Disch pressure Transducer failure Installation

D030-204 LH2 Recirc Pump Disch Pressure Transducerfailure Installation

DO30-205 LH2 Recirc Pump Disch Pressure Transducer failure Installation

D152-202 LH2 Recirc Pump Inlet Pressure Transducer failure InstallationS013-2i8 Stringer 20 Side Long Strain Transducer open

SIVBSTAGE I

I_254-403 Press-LOX Tank Repress Sphere Drifted low

LOOIg-4O8 Level-LiquidHydrogen Pos C Dropped out when wet Return to "ON"state after

several minutes.

I NS TR UM EN T UN IT

Dl7-6Ol Methanol/tJaterCoolant Pump Exit Noisy iJaivedduring CDDT. Provided

pressure useful data during flight.

19-2



"F able 19 -3. A S -50 5 M easurement M alfunctions

TIM[ OF

MEASUREMENT MEASDR[HENT TIILE UAFIIR[ OF FAILURE FAILIPRF gURATIOI4NUNgER (I(ANG[ SATISFACTORY RFiIARKS

TIME) OPERAIIOM

TOTAL HEASUREI'IENTFAILURFS_ SoIC STAGE

BgI5-118 Acoustic, External Data low 0 sec Done

C033_I06 Heat Shield, Internal Transducer failure O sec None DO usable data.

0547-106 Differential, Heat Shield Protective cover for , 0 sec _lone Positive port sealed.i sen _in g p o rt n o t rem oved

i b efore flight

E038-101 Vibration, Fuel Pump ; Data level too high Ignition gone

Flange, Rad

EOSd-115 Vibration, Ret_raotor Trahsducer failure. See Ren_rks. _one NO data after telemetry

$witchover.

TOT AL M EA SU RE IIE gT F AIL UR ES, S -I I STA GE

DllB-21g Fo_ard Skirt Static P Transducer failure. 0 sec )Done

EOOI-203 E3 Long Vib Conbstn Dome Cable and/or connector O sec None

open.

EODB-204 E4 Radial Vib Fuel Pu_) Cable and/or connector 0 see _one

open.

E326-219 Long gib Pwd Skirt Cab)e and/or connector 0 sec 'lone

Stringer open.

E339-Bg6 Nom Vib Thrust Cone Cable and/or connector 0 sec 'loneOpen.

E352-206 Tan Vib lnters tage Frame Cable and/or connector 0 sec _one

open.

TOTAL MEASUREMENTFA[LURES_ S-[VB STAGE, PHASE I ALSO INCLUOED IN PHASE II)

BOOlg-427 Acous Aft Skirt Envelope decrease. 65 sec 65 sec Decrease in amplifier

SEa 2880 Ext _gain suspected.

50025-426 Acous SEa 32BB, Pos I Envelope decrease, 45 sec 45 sec Decrease in mplifler

Ext Bin suspected.

00200-401 Temp-Fuel Injection 6 to 8 percent low -I0 minutes _tone Unknown.

00230-403 Press-GBX/GB 2 Burner lO percent upward shift Prior to _ione Sensor problem.

GH2 [nj ltftoff

ANBITI_At MEAS_gEHE_T FAILDRES, S-IgB STAG_, P_ASE IF

A0010-403" Accel-Bimbal Data drifted to upper 1200 sec 1200 sec Low temperature failure.BIock-Pitch-LF band edge.

bBl04-4O3 P_ess-LH B Press Module Steady state varied 9270 sec 9270 sec Degradation of amplifier

Inlet during second burn suspected.

D0236-403 P_ess-Ambie_t Uelium Off scale high at 975q gTso sec g75Q sec Posslble oDen bridge.

Pneu Sphere seconds.

E0239-401 Vib-LOX Turbine Bypass Data erratic during 9210 sec gBlO sec Coaxial cable

Vlv Tan second engine burn discontinuity is

suspected.

PARTIAL MEASUREtlENT FAILURES, S-IC STAGE

A001-118 Accel, Long Data went to negative O seca t l if to ff

BOO4-114 Acoustic, Fin D Nigh amplitude low O sec

f re qu en cy n oi se

B006-118 Acoustic, External Bias leve) too high ll5 sec Positive clipping and

negative

BOO7-118 Acoustic, External Bias level too high 37 see Positive clippil}g

BDOD-IID Acoustic, External Bias level too high 51 to El sec Positive clipping

I)hF)%IIB Acoustic, [xternal Data clipping 37 sec Positive clipping

BF)IO-II_ Acoustic_ [xternal Data clipping lID sec Positive clippit_g

DOll-liFt Acoustic_ [_ternal Data clipping 0 sec Some u_ahle data.

[;61]B_ll_: AcoLJ_tic, FxternaI Data Clippin(l 0 _P£ _ol)leu_able data.

li(iIl-I1:_ AcoLJsLic, Ixternal Data cli_pill_[ O s_c So_(,i_s,_bledata.

*Cm_tractor position is that this ll_asurei_!ntfailed o_tside the period o( iiItel-esL.

19-3



T able 1 9 -3. A S -50 5 M easurement M alfunctions (Continued)

T IM E O FMEASUREMENT fAILURE DURATION

NUMBER NEASUREHENFTITLE NATUREOF FAILURE (RANGE 5ATISFACTDR_ REIIARKSTIME) OPERATION

PARTIALMEASUREMENTFAILURES,S-iCSTAGE(Continued

BOIR-]I8 Acoustic,External Data clipping 0 sec IntermittentSome usabledata.

COO3-1O2 Temp.TurbineManlfoid Data dropsunexpectedly 115 sec llS sec

C131-105 TemperatureSolenoid Data noisy 35 sec 35 sec Minor lossof data.Valve

DOII-IOI Press,Dng Control Data leveltoo hlgh 70 sec 10 sec

0020-101 HeatExchangerOutlet Transducerfallure 52 sec i2 sec Data usablepriQrt_52 sec.

D145-]i5 Helium Inlet Data trendIC_ 7 sec sec Data usable.

0150-I15 LOX PumpInlet Transducerfailure I00 sec IO0 se¢ Data usableprior to100sec.

E042-102 FuelPump Flange Radial High amplitudelow 0 sec [ntennltteniSome usabledata.frequency noise

E042-I03 FuelPumpFlangeRadial Highamplitudelow 0 sec Intermittentomeusabledata.

_eq_eBcy noise

E042-I04 Vibration,Fuel p_np Data dropouts 0 sec IntemlttentFl ange R adial

D007-I02 PressBre,Fuel Pump Unexpecteddecreasefrom 85 sec All but TODischarge,Engine2 85 to 95 sec seconds

D007-I04 Pressure,Fuel Pump Unexpecteddecreasefrom 50 sec All but I0Discharge,Engine4 50 to 60 sec seconds

PARTIAL_EASURE_EIITAILURES,S-II STAGE

C003-201 El Fuel TurbineInletT Transducerfailure 177.5sec 177.5sec

C680-206 HeatShieldAft Gas T Intermittentand noisythroughout

C72T-206 Heat ShieldAftHeat Rate Transducerfailure 195 sec 195sec

C815-206 LOX Vent ValveMeat Bate Transducerfailure 102 sec I02 set

D012-202 E2 EngineReg OutletP Transducerfailure 210 sec _I0 sec

DO6D-2OO UllageRocket8 ChamberP 10 percentOC bias shiftin transducer

0100-206 Heat ShieldFwd FaceP Transducerfailure 200 sec 200 sec

EDOI-204 E4 Long Vib CombsbnDc_ne Periodsof amplifiersaturation

E002-203 E3RadialVibLOXPump Periodsofamplifier

saturation

E002-204 E4 Radial Vlb LOX Pump PeriodsoF amplifiersaturation

E003-203 E3RadialVib FuelPump Periodsofamplifier

saturation

E215-202 E5 Rad Main Fuel Valve Amplifiersaturation 26l to 276 _537secset

E342-203 Tan Vib LH2 Prevlv/Fdln Periodsof amplifiersaturation

PARTIALMEASUREMENTFAILURES,S-IVBSTAGE

BOOl6-Rll AcousticFwd Skirt ApproximatelyIg data 65 sec All but 20 A loosecable is

Station3216-iMT dropoutswere observed sec suspected.between 65 and 85 sec

CODOI-40i Temp-FuelTurbineinlet Data lowerthan expected 0 sec Data believedto beusable.

COigg-4ol Temp-ThrustCham_er SIO_ responseduringFirst firstburn Slightsensocdebondteg

Jacket engine burn duringFirstengineburnis suspected.

O0233-403 Press-O2H InjSpool Approximately8 percent O to 20 sec All bud 20 SusceptibletoChamberO_f noise fromlifteffto seconds vibration.

liftoffplus20 seconds.

Eoog2-4OR Vibration-Station748 Data level lowerthan 0 sec Data believedto be

PositionII-Thrust expected usable.

19-4



Table 19-4. AS-505 Questionable Flight Measurements

R E A S O N

MEASUREMENT MEASUREMENTITLE QUESTIONED REMARKSNUMBER

QUE ST IO NA BL E M EA SUR E ME NT S, S -II S TA GE

DI16-219 ForwardSkirtStaticP Decaynot proper; Full scalerangeof instrumentquestionableafter isextremelylow.50 seconds.

DI17-219 ForwardSkirtStaticP Decaynotproper; Fullscalerangeof instrument

questionable after is extremely low,50 seconds.

D120-219 ForwardSkirt Static P Decay not a s predicted. Full scale range of instrumenti s e xt re me ly l ow.

D121-219 ForwardSkirtStaticP Decaynot proper; Fullscalerangeof instrumentquestionable after is extremely low.

_75 seconds.

19.3 AIRBORNE TELEMETRY SYSTEMS

Performance of the 17 VHF telemetry links was generally satisfactory with

the minor exceptions noted. A brief performance summary of these links

is shown in Table 19-5.

There was a variation of approximately lO counts in the lO0 percent level

of the inflight calibrations for the DP-I telemetry link. This is equiv-

alent to 50 millivolts as compared to 41 millivolts in the specifications.

This type of high variation has been previously experienced on AS-205 and

during checkout. E xamination of 5-vdc measuring voltage supply data asseen on word 28-frame lO of the DPI-AO links also indicates variations of

this magnitude. This problem is under further investigation.

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 19.5.1.

Usable VHF telemetry data were received to 15,780 seconds (04:23:00) at

both GYM and Hawaii (HAW).

Performance of the CCS telemetry was generally satisfactory except for the

period during translunar coast from 23,601 seconds (06:33:21) to 25,111seconds ( 06:58:17). T his problem is discussed in detail in paragraph

19.5.3.2. Usable CCS data were received at GDS to 39,305 seconds (I0:55:05).

19.4 AIRBORNE TAPE RECORDERS

The performance of the three onboard tape recorders installed to record

real time data during predicted R F blackout periods was satisfactory.

Noise levels, timer operations and recorder response times remained within

19-5





required operating limits. Recorded data agreed favorably with data

obtained in real-time. Approximately 51.1 seconds of S-IC data and179.7 seconds of S-II data were recorded. All of the recorded data were

successfullyplayedback.

Recorder assignments and their periods of performance are listed in

T able 19-6.

19.5 RF SYSTEMS EVALUATION

The performance of the RF systems, based on data received to date, was

good. Measured Flight data, with few exceptions agreed favorably with

expected trends. RF performance of the telemetry, SR SCS and tracking

systems was good. CCS performance was generally satisfactory with the

exception of the problem discussed in paragraph 19.5.3.

VHF final LOS was reported by BDA at 18,900 seconds (05:15:00) and CCS

LOS at 40,191 seconds (II:09:51) by GYM and GDS. BDA indicated C-Band

tracking L OS at 35,346 second (09:49:06) .

19.5.1 Telemet[y System RF Propagation Evaluation

The performance of the 17 VHF telemetry links was excellent and generally

agreed with predictions. Ultra High F requency (UHF ) telemetry link DP-IAwas deleted on AS-505.

Moderate to severesignalattenuationwas experiencedat varioustimes

duringthe boost due to main engineflameeffects,S-IC/S-IIand S-II/S-IVB

staging, S-II engine ignition and S-II second plane separation. Magnitude

of these effects was comparable to that experienced on previous flights.

At S-IC/S-IIstaging,signalstrengthon all VHF telemetrylinksand on

the CCS downlink dropped to " threshold for approximately I and 5.6 secondsrespectively. Signal degradation due to S -II engine 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 much of the data was recovered from

onboard tape recorder playback, other stations providing overlapping

coverage, or 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, although the Mercury (MER) ship experienced a drop in RF signal

strength to -127 dbm, shortly after start of S-IVB R estart Preparations

(Time Base 6). This dropout was at least 90 seconds in duration. Valid

data were received during this period from Carnarvon (C RO ), indicating that

vehicleinstrumentationsystemswere operatingsatisfactorily. Performancewas nominal during second burn and final coast, except for the CCS downlink

problem discussed in paragraph 19.5.3.2.

1 9 - 7



m VAN

PARKINGORBIT It4SERT_ONmmmmmmlBOA BEGINS-[VB RESTARTPREPARATIONS_S-[VB SECOND IGNITION

GBM _ TRANSLUNARINJECTION

GBI

_ER

..mm. MILA _I_ CYI _ CRO _m_GYM

I,_o _2_o _Coo 2_o 30'0o _o 4_Qo _ s4'oo

RANGE TIME, SECONDS

' '_ ' ' I , l 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 O1:30:00

RANGE TIME, H0oRS:MINb'TES:SEGOI_DS

m BOA

m GBN

i II_ MILA m RED

i m TEX mlmmm CYI m MER

GYM_ VAN g_ CRO

i s4oo 0o_o 6doo 72;0 Woo _oo o&o o6oo' lO,'2oo io:sooRANGETIME, SECONDS

,-, , , , , , I0 21 _{_ I _' _', l I i1:30:00 01:_,0:00 Ol:SO:O0 0o:00:00 02:1 :00 02:0:00 02:30:00 02:40:00 02:50:00 03:00:00

RANGE TIME, HOURS:MINUTES:SECONOS

BDA

m TEX

G YM

l I HAW

7200 m,_oo 14,:=oo _8,;oo 21,;oo zs,)ooRANGE TIME, SECONOS

I, . ,,I , I , I I I

02:00:00 03:0O:00 04:00:0O OS:0O:O0 06:00:00 07:00:00


F igure 19-1 . VH F Telemetry C overage Summary

BDA reportedVHF LOS at 18,900seconds (05:15:00)and GYM and GDS reportedCCSLOSat 40,191 seconds (11:09:51).

A summary of available VHF telemetry coverage showing Acquisition ofSignal(AOS)and LOS for each stationis shown in Figure19-I.

19.5.2 Tracking Systems RF Propagation Evaluation

Analysisof data receivedto date indicatesthat the C-Band radar functioned

satisfactorily during this flight, although several stations experiencedtracking problems due to phase front distortions and equipment malfunctions.

The ODOP system,previouslyflownon the S-IC stage,was deletedon thisflight.

19-8





_PARKING ORBIT INSERTION

mlmmBDk FPQ-6 _/DEGIN S-IVBRESTART PREPARATIONS

BOA FPS-16 _S-IVB SECOND IGNITION

_GTI _'TRANSLUNAR INJECT[Obl

i GBI

i PAFB

aim CNV _mmmm MER

mmmim MILA _ITAN i CRO

6_Ov ' J i200 1_00 2_00 3doo 3_00 4Joo 4o00 siooR AN GE T IME , S EC ON DS

t I _ # / I I I I00:00:00 O0:IO:O0 00:20:00 00:30:00 010:40:00 00:50:00 01:00:00 :_0:00 01:210:00 01:310:00


I i BOA FPQ-6

i BOA FPS-16

GTI

PAFB mmi HER

i MILA _II TAN m CRO HAW fi i !

5400 GJoo 66'00 7/00 Hoo 84'00 Rdoo 0600' 10,200i0,800RANGE TIME, SECONDS

Ol:3(]:00 ON:40:00 ON:50:00 02:O0:00 02:10:00 02:20:00 02:30:00 02:40:00 02:50:00 03:00:00


m ANT

PAFB

II GIN

BOA FPQ-Bl HAW

! ! i i ! i m ! ! m

7200 10,000 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 | I2:00:00 03:00:00 04:00:00 05:00:00 06: :00 07:00:00 08:00:00 09:00:00 10:00:00 I1:00:00


Figure 19-2. C-Band Radar Coverage Summary

19.5.3.2 Command and Communications System. Available data indicated

satisfactory CCS performance during boost and parking orbit with minor

exceptions. Downlinkdropoutsoccurredduring S-IC/S-IIstaging and atS-II second plane separation. Dropouts at these times are expected.

S tation handover was accomplished with minimum data loss (less than 5

seconds) from MILA to BOA at 362 seconds and from BDA to VAN at 690 seconds.

Downlink dropouts were also experienced during the second revolution during

handoverfromMILA toGran

dBahama (GBM) at 5880.5seconds (Ol:36:00.5)

and during handover from GBM to BDA at 6120.3 seconds (01:42:00.3).

Duration of these dropouts were 9.5 and 5.2 seconds respectively.

Performance during second burn and translunar injection was nominal.

19-10



During the final coast, a sharp drop in downlink CCS signal strength wasnoted at GYM and GDS at 23,601 seconds (06:33:21). T he onboard antenna

system, which had been on low gain since 15,233 seconds (04:13:53), was

switched to the high gain mode at 24,052 seconds (06:40:52) to improve

signal quality. Signal strength picked up and was maintained at a high

level until 24,160 seconds ( 06:42:40) at which time the signal level again

dropped, then was completely lost approximately l minute later. Thesignal fluctuated intermittently at low levels until 25,097 seconds

(06:58:17) at GDS and 25,111 seconds (06:58:31) at GYM. The system, whichhad been commanded to the omni-directional mode at 25,097 seconds

(06:58:17), remained in this mode until final loss of signal at 40,191

seconds (II:09:51). F igure 19-3 shows the fluctuation in signal level

experienced during this time at the GYM site. The GDS station experiencedsimilar fluctuations at corresponding times as shown in Figure 19-4. Signal

levels were slightly higher at GDS due to the 85 foot antenna used versusthe 30 foot antenna at G Y M .

N ormally, the directional high and low gain antennas would be expected to

provide higher signal levels than the omni-directional antennas. Duringthis time period the vehicle was at a sufficiently high elevation angleand vehicle attitude was such that good signal could be expected,

- 1 0 0 ,

-110 •

-120.

=

-I 40,

-150CCS]OMNLIhlKIGNAL

(30 ft DISH)

23,400 23,580 23,760 23,940 24,120 24,300 24,480 24,660 24,840 25,020

RANGEIME, SECONDS

I ! I I I I I I I I

06:30:00 06:33:00 06:36:00 06:39:00 06:42:00 06:45:00 06:48:00 06:51:00 06:54:D0 06:57:00

RANGEIME, HOURS:MINUTES:SECONDS

Figure 19-3. CCS Signal Strength F luctuations at Guaymas

1 9 - 1 1



p,o[DS

-80 -- C----CSOWRLINR SIGNAL___ SIGNAL BECOMESSTEADY

(85 ft. DISH) --'_25,097 SECOHDS (06:58:17

T O HIGH GA IN

- 9 0

-lO0

-]lO

-140

-160

-17023,400 23,580 23,760 23,940 24,120 24,300 24,480 24,660 24,840 25,020

RANGETIME, SECONDS

|, 1 I I I i I I I I

06:30:00 06:33:00 06:36:00 06:39:00 06:42:00 06:45:00 06:48:00 06:51:00 06:54:00 06:57:00


Figure 19-4. CCSSignal Strength Fluctuations at Goldstone

1 9 - 1 2





CCS

RCV

lu cI-CCsRANSPONDER ANTENNA

DECODER I .

I 1 .J4 LOWGAIN(70)

:_ l z2_2.sz'__ ccs.,TDIRECTIONAL

1 ._] COAXIAL J2 ANTENNA

-- SWITCH

SW SEL XMTRDISABLE

RF CCS/PCM

OMNI #I

IDIVIDERICS'"C"OMNI 02

F igure 19 -5 . C C S System Block D iagram

D TK I H- IO - 6P

J 5

+28VDC B ..I

!REL_A i _C_ICOM

i J 4

--I

I

F __I IND CKT

RF POS J3 D___IIND CKT

L COM

"_"_j_ _ ? E INDKT__2

AJ2

+28 VDC

Figure 19-6. Electrical Schematic of CCS Coaxial Switch

19-14



_ TP AR KIN G O RB IT IN SE RT IO N

_B EGIN S-IVB RE STA RT PR EPA RAT ION

1CYI _/S-IVB SECOND I_ITION

IB VAN _TRANSLUNARNJECTION

HSK GDS (

lll MIL llCRO GYM_r

6"oo ' ' o" L ' ' ' '200 1800 24 0 30uO 5400600 4200 4800


# 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 0i:00:00 Ol:IIO:O0 01:20:00 01:30:00

RA N GE T IME, HOU RS:MI NUT E S :S E C ON D S

IGDS IGBM

NmVAN

lml,m BDAIRED

mmmmmmi, MIL

llTEXCYI n CRO

1GYM

5400 60_0 6600n 7_00 7,_0 8_00 go00' 9600 10,200' '0,800

RANGE TIME, SECONDS

, , _ ,_ _, ,Ol:30:OO 01:40:00 Ol:SO:O0 i ' 02:30:00 03:00:002:00:00 02:110:00 02:20:00 02:40:00 02:50:00

RA N GE T IM E, HOURS :MI NUT E S :S E C ON D S

FINAL LOSGDS

GYP

! I ! I I f I I I

7200 10,800 14,400 18,000 21,600 25,200 28,80D 32,400 36,000 39,600

RANGE TIME, SECONDS

' ' 010 I I 0 I I oio I L2:00:00 03:00:00 04: :00 05:00:00 06:00:00 07:0:00 08:00:00 09: :O0 10:00:00 11:00:00

RANGE TIME, HOURS:MINUTES:SECOflDS

Figure 19-7. CCS Coverage Summary

19-15/19-16



S ECTIO N 20

M A S S C HA R A CT E R IS T IC S

20 .1 S UM M A R Y

P ostflight analysis indicates that tota l vehicle mass was with in 0 .50

percent of the prediction from ground ign ition through S -IVB stage finalshutdown. T his very small deviation signifies that the initia l propellantloads and propellant utilization throughout vehicle operation were close

to predicted.

20 .2 M AS S E VA LUA T IO N

Postflightmass characteristicsare comparedwith the finalpredictedmasscharacteristics ( MS FC Memorandum S &E -A ST N-S AE -69-M-53) which were used in

the determination of the final operational trajectory (MSFC Memorandum

S&E-AERO-FMT-106-69).

The postflight mass characteristics were determined from an analysis ofall available actual and reconstructed data from S-IC stage ignition

through S-IVB stage second burn cutoff. Dry weights of the launch vehiclewere based on actual stage weighings and evaluation of the weight and

balance log books (MSFCForm 998). Propellant loading and utilization wasevaluated from propulsion system performance reconstructions. Spacecraftdata were obtained from the MannedSpacecraft Center (MSC).

Deviations from predicted in dry weights of the inert stages and the loadedspacecraft were all less than 0 .5 percent which is well within the 3-sigm adeviation limits.

During S -IC powered flight, m ass of the total vehicle was determined tobe 26 7 3 kilogram s (589 2 Ibm ) or 0 .0 9 percent lower than predicted at

ignition, and 2125 kilograms (4684 Ibm) or 0.25 percent lower at S-IC/S-IIseparation. T hese very small devia tions are attributed to a less than

predicted S -IC propellant 'load and a slightly less than predicted upperstage m ass. S -IC burn phase tota l vehicle m ass is shown in T ables 20 -Iand 20-2.

During S-ll burn phase, the total vehicle mass varied from 142 kilograms(313 Ibm) or 0.02 percent lower than predicted at start commandto 559kilograms (1233 Ibm) or 0.27 percent higher than predicted at S-II/S-IVBseparation. The initial deviation may be attributed to a slightly less

2 0 - I



than predicted tota l propellant loading, and the deviation at separationwas due m ain ly to a higher than predicted S -II L O X residual. T ota lvehicle m ass for the S -II burn phase is shown in T ables 20 -3 and 20 -4.

During S -IVB stage operation, the total vehicle mass varied from 30 9

kilograms(681 Ibm) or 0,19 percenthigherthan predictedat firststart

command to 260 kilograms (572 Ibm) or 0.42 percent higher than predictedat end of second burn thrust decay. These deviations are due mainly to

a slight excess of S-IVB propellants. Total vehicle mass at spacecraft

separation was 367 kilograms (808 Ibm) or 2.04 percent higher than

predicted. Tables 20-5 through 20-8 show the vehicle mass history during

both S-IVB burn phases.

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 gravityand moment of inertia is shown in Table 20-I0.

20-2



Table 20-I. Total Vehicle Mass - S-IC Burn Phase - Kilograms

6ROUND IGNITION HOLDOOWN CENTER OUTBOARD S-IC/S-II

EVENTS ARM RELEASE EN@INE CUTOFF ENGINE C_TOFF SEPARATION

PRED ACT PRED ACT ORED ACT PREO ACT PRED ACT

RANGE TIME--SEC -6,39 -6.37 ,25 ,25 135o23 135o16. 160,20 161.G3 163,91 1_2,31

S-IC STAGE DRY 1334_7, 1333_, 133_7. 133344o 133447. 133344o 133_47, 1333_4. 133_47° 13334_.

LOX IN TANK Iq77931. 1477050. 1446718° Iq43231. 189q09° 19q713. 668. 1559. 599° 1043.

LOX _ELOW TANK 21000° 21087. 21737. 21847° 21720. 21778. 16761. 16853. 14629. 14994.

LOX ULLAGE BAG 190° 191. 210. 222. 2590° 2878. 3062. 3377. 3068. 3382°

RPI IN TANK 6_Z8BZ. 641271° 6327_9, 630605° gZ682o _1987. 9997° 6993. 8697. 5B_.

RP1 BELOW TANK 4313, q306, 5996. 5989. 5996° 5989° 5958. 5951. 5958° 5951.

RP1 ULLAGE _AS 35. 73. 35, 76, 211o 230, 240. 254. 241° 255.

N2 PURGE GAS 36. 36° 36° 36. 20° 20° 20. 20. 20. 20°

Po HELIUM IN BOTTLE 289. 28_° 289. 285. 113. 132. B3. 108. 83. I07.

0 FROST 635. 635° 635. 635. 3_0. 3qO. 340. 340° 3wO 340.I °

C_ RETROMOTOR PROP i027° IO27. ID27. 1027° 1027o I027° 1027. 1027° ID27° 10_7°

OTHER 239. 239° 239. 239° 239. 239° 239° 239. 239. 239.

TOTAL S-IC STAGE 2282034° 2279549. 2243118. 2237537. 4q7793. 452677. 171842. 170065. 168548. 166566.

TOTAL S-IC/S-II IS 5262. 5255° 5262° 5255. 5262. 5Z55. 5262° 5255° 5229. 5222.

TOTAL S-II STAGE 484590. 454159. 48qSBG° W84159. _84331o 483901. 4Bq331, 483901° _8_331° 483901°

TOT S-II/S-IVB IS 3665. 36W9° 3665. 3649° 3665. 36_9° '3665. 3649. 3665. 36_9.

TOTAL S-IVB STAGE 118858, 119223° llBB58, 119223. 118722, 119132° 118722, 119132° llaTZ_. 119137°

TOTAL INSTRU UNIT 1930o 1935. 193D° I93S° 1930. 1935° 1930° 1935. 193D° 1935.

TOTAL SPACECRAFT _8731° 48625. 48731, 48525. 48731. 4BG25. 48731° 48625. _8731° 48625.

TOTAL UPPER _TAGE 663035, 6628_7. 663035° 6_2_7. 662641° 662_97. 662641° B62497o 662607. G62W6_,

TOTAL VEHICLE 2945069. 2Bq2396° 2906153° 2900383° 111_43_o 1115175. _3q_B3° 832562, 831155. 829D31°



T able 20 -2. T o ta l V eh icle M ass - S -IC B urn P hase - P ounds M ass

GROUND IGNITION HOLDOOWN CENTER OUTBOARD S-IC/S-II

EVENTS ARM RELEASE ENGINE CUTOFF ENGINE CUTOFF SEPARATION

PRED ACT PRED ACT PRED ACT PRED ACT PREO ACT

RANGE TIME--SEC -6.39 -6.37 .25 .25 135.26 135.16 160.20 161.63 160.91 162.31

i

5-IC STAGE DRY 294200. 293974. 29420G. 293974. 294200. 293974. 294200. 293974. 294?00. 293974.

LOX IN TANK 3258280. 3256338. 315948_. 318178D, W17575. 4292_8. 1472. 3437. 1321. 2303.

LOX BELOW TANK 46296. 46489. 47921° 48164° 47884. 48013. 36953. 37155. 32251. 33056.

LOX ULLAGE SAS q18. 422. 463. 489. 5710, 6346. 6750. 7444. 6763. 7457.

RP1 IN TANK 1417335. 1413761. 139_972. 1390245. 204328. 202797. 22039. 15417. 19615° 12927.

RPI BELOW TANK 9509. 9493° 13219° 132D3, 13219. 132D3° 13136. 1312D° 13136. 13123.

RP1 ULLAGE GAS 77. 161. 77, 168. 464. 507. 529. 560. 531. 562.

N2 PURGE GAS 80. 80. 8D. 80. 43. 43. 43. 43. 43. 43.

HELIUM IN @OTTLE 636. 636. E36. 629. 249. 290. 183. 237° 182. 235.r_0 FROST 1400. 1400. 1400. 1400. 750. 750. 750. 750. 750. 753.

I RETROMOTOR PROP 2264. 2264. 2264. 2264. 2264. 2264 2264. 2264. 2264. 2264..OTH[R 528° 528° 528. 528. 528. 528. 528. 528, 528° 52B°

TOTAL S-IC STAGE 5D31024. 5825546. 4945228. 4932924° 98#215. 997983. 378848. 374929. 371584, 36721_..............................................................................................................

TOTAL S-IC/S-II IS 11600. 11585. 11600. 11585. ll&OD. 11585. 116DO. 11585. 11527. 11512.

TOTAL S-II $TA6£ 1C68337, 10E7389. 1068_37. 1067389. 1067767. 1066819. 1067767. 1D66819, ID67767. 1D66819°TOT S-II/5-1VB IS 6081. 8045. 8G81. 8045. 8081. 8045. 8081. 8045. 8381. 8045.

TOTAL S-IVB STAGE 262037. 2628_1. 262037. 262841. 261737. 262641. 261737. 2626_1. 261737. 2626_1.

TOTAL INSTRU UNIT _254° 4267. 4254° 4267. 4254. 4267. 4_54. 42_7. 4254. 42_7.

TOTAL SPACECRAFT I_7433. 107200. 107433. 107200. 107433. 107200. 10743_. 107200. 107433° 107200..............................................................................................................

TOTAL UPPER STAGE 1461742. 1461327. 1461742. 1461327. 146C872° 1460557. 146D872. 1460557. 1460799. 1460484.

TOTAL VEHICLE 6492766. 6486873. 6406970. 6394251. 2_48087. 2458540. 1839720. 1835_86. 1832384. 1827700.





Table 20-4. Total Vehicle Mass - S-II Burn Phase - Pounds Mass

S-IC IGNITION S-!I S-II S-IT S-II/S-IV_

EVENTS IGNITION MAINSIAGE ENGINE CUTOFF SEPARATION

PRED ACT PRED ACT PRED ACT PRED ACT PRED ACT

RANGE T!ME--SEC -6.39 -6.37 162.61 16W.D5 ISW.SI 166o30 SSW°I3 552.64 555oD4 553.50

G-!C/S-!! IS SMALL 1350° 1348°

S-lC/S-II IS LAR3E 8890. 8892. 8590. 8892. 8890. 8892.

S-IC/S-II IS PROP 13GO. 1345. 689. 682° 0o O°............................................................................... ..............................

TOTAL S-IC/S-II IS 116C0. 11585. 9579. 9574. 8890. 8892.

S-II STAGE DRY @4367. 84273. 84367, 84273. 84367° 84273° 84367. e4273. 84367. 84273,

LOX IN TANK 82287_° S21700. 82297_. 8217D0° 821856. 920714° 1447. IaO0° 1197o 1513,

r,_ LOX BELOW TANK 1625. 1625. 1625. 1625, 1764. 1764. 1736. 1736, 1736. 1736.0 LOX ULLAGE _AS 395. W05° 3_5° 405. 399. 409. 5114° 5387. 5119. 53_7°

I

LH2 IN TANK 158000. 158310. 157986. 158295, 157516. 157834. 4335. 4350, 4224, 4254°

LH2 bELOW TANK 231° 231° Z45o 246. 282° 282° 272. 272. Z72. 272.

LH2 ULLAGE GAS 169° 169. 1&9° 169, 171. 171° 1561. 1625° 1562, 1625°

INSULATION PURBE 120. 120.

FROST 450. 450.

START TANK GAS 30. 30. 3D. 30. 5. 5. 5. S. 5. 5.

OTHER 76. 76, 76. 7&. 76. 78. 76. 76. 76. 76,

.................................... _ .................................... ....................................

TOTAL S-II STAGE 1068337. I0673B9. 1067767. I_86819° 10E6436° lOG5S2e. 98913. _9524. 98558. 99146.

TOT S-!IIS-!_E I_ 8081. 80_5. 8081. 80_5. &081. BOWS. 8081. 80_5. 8581. BOWS.

TOTAL S-IVg STA_E 262037, 2_2841° 251737. 2_2841° 281737. 2&2641. 2G1737. 2_2541° 261732, _8_B3_.

TOTAL INSTRU UNIT 4254. 4267° 475_° 4267° 4254. 4267. _254° _267. 4254. 4267.

TOTAL S_ACZCRA_I I0743_. 107200. 1C7_33o 107200. I87433. 10720D. 9_500. q_264o 98530° 98254.

TeTAL UPPER STAGE 381885° 382393. 381505. 382153. 381505. 352153. 372572° 373217° 372567, 373212.

TOTAL VEHICLE 1461742° 1461327° 1458851. 1458546. 1496831. I_56573. _71_85. 4727wi, 471125° 472358.



Table 20-5. Total Vehicle Mass - S-IVB First Burn Phase - Kilograms

S-IC IGNITION S-IV8 $-IVB S-IVB S-IVB

EVENTS _GNITION MAINSTAGE ENGINECUTOFF Eh:ODECAY

P_ED ACT PRED ACT PRED ACT P_EO ACT PRED ACT

RANGE TIME--SEC -6.39 -6.37 558.14 556.81 560.6_ 559.31 703.54 703.76 703°74 703.98

S-IVB STAGE DRY I168_. 11648. 11657. 11525. 11657. 11625. 11596. 11563. I1596° I1563.

LOX IN TANK Bb53@. R696_. 96539. 8696_o 86376, B6831+ G030G. 6QS4Bo 63275, 9351_.

LCX BELOW TANK 166. 166. 166° 166. 18D. 180, 190o 190. 180. 180°

LOX ULLA3Z GAS 18. 22. 18. 22. 25. 23. 102. 100, 102. 131.

LH2 IN TANK 197_9. 19659. 19705. 1965W. 196_9. 19598. 14429. IW291. 14W15o IW277.

LH2 BELOW TANK 22. 22. 26, 26. 2&. 26, 26. 2_. 26+ _&+

LH2 ULLAGE GAS 20. 21. 2D. 21. 21. 21. 66° 65. 67. 66.

ULLAGE_OTOR PRO_ 54. 54. iD. i0. I. I. I. I. I, I,

AP$ PROPELLANT 286. 3C3. 286. 303. 286. 363. 295. 298. 285. 298.

_ELIUM IN 50TTLZS 20C. 200. 200. 200. 200, 200° 179, 176. 17_o I?6.rO

5TARTTANKGAS 2. 2, 2. 2. O, 0. 3o 3o 3° 3.

l FROST 136. I3G. 0. 45° O. US, O. _5+ 0. _5°

OTHER _5, 25, 25. 25. 25. 25° 25. 25. 25. 25.

TCTAL S-IV_ STAGE 118858. 119223. 118655. I19065° 11_4W5o 11B84_o Q7197° e7322° 87152. 87278.

TO_AL I_ISTFU UNIT 1930. 1935. 1930. 1935. 1930. 1935. 1930. 1935. 1930. 1935................................................ ................................... ...........................

TOTAL SPACECRAFT W_679. W4572. 44_79. 4W572, W_679. 44572. 4_679. W_572. 4_679. W4572........................................... ......................... ..........................................

TOTAL UPPE p STAGE q6608. 46507. 46_8. 4_507. 46608° _6507. W6608o QG_DT° _6608. q65C7°

TOTAL VEHICLE 16546_o 165730. 165263. 165573. 16505ho 165355° 133806. 133_30. 133760. 133786,



Table 20-6. Total Vehicle Mass - S-IVB First Burn Phase - Pounds Mass

_-IC IGNITION 5-1VB S-IVB S-IVB S-IVB

EVENTS IGNITION HAIN_TA3E ENSINE CUTOFF END DECAY

PRED ACT PRED ACT PREO ACT PREO ACT PRED ACT

RANGE TIMZ--S_C -6.39 -6.37 556.I; 556,$1 560,64 559o31 703.54 703.76 733.74 733.9_

S-IV5 STAGE DRY 2575C. 25680° 25699. 25629. 25699. 25626. 25_64° 25;92. 25564. 25492.

LOX IN TANK 190765° 191722. 190785, 191722. 190426, 191363. 132953° 133;86. 132683° 133416,

LOX BELOW TANK 367. 367. 367. 367. 397. 397, 397. 397. 397. 397.

LOX ULLAGE GAS #0, 49. 40. 49° 56. 50, 224. 221° 225, 223°

LH2 IN TAN_ ;3;52o ;3340, q3_42. 43330. 43318. 4320B. 31810. 315Q6° 31779° 31476.

LH? 3ELOW TANK qS. 48. 58. 58. 58. 58. 58. 58. 5R. 56.

LH2 ULLAGE _AS 45, 46. 45. 46. 46. 46. 147. I44. 147. 145.

ULLAGE MOTOR PROP 118. 118. 22. 22. 0. O. O. 0. O. O.

APS oROPELLANT 630. 668. &30° BGBo 630. 668. 628. 65_. 629. 656.r,30 HELIUM It; eOTTLE_ 441. 442. 441, 442. q;O. 441. 393. 388, 392. 388,

Oo START TANK 3AS 5. 5. 5o 5. i. I. 7. 7. 7. 7.FROST 300. 300. O. I0_. 0. IOC. G. 100. O. lOO.

OTHEQ 56. 56, 56. 56. 5_. 58. 56. 56. 56, 5&.

TOTAL 5-IV8 STAG_ 262037, 262841. 261589. 262494° 261127. 262015. 192237° 192513° 192137. 19241_o......................................... ....................................................................

TOTAL I_STRU UNIT 425;. 4267° 4254° 4267. 4254. 4267. 4_54° 4267. 4254. _267.

TOTAL SPACECRAFT 9850_. 98264. 98500. 98264. 9550D. 68264. 9_500. 9_264. _6500. 9526_.

TOTAL UPPE_ 5TAGC i02754. 102531° 102754, 102531. 102754. 1D2531. 102754. 102531. 13_754. ID2531.

TOTAL VEHICLE 364791. 365372° 364343. 365025° 363881, 3_45q6. 294391° 295044. 294891, 29_94T°



T able 20 -7 . T ota l V ehic le M ass - S -IVB S econd B urn P hase - Kilograms

S-IVB S-IVB S-IVB S-IVB SPACECRAFT

EVENTS IGNITION HAINSTASE ENGINE CUTOFF END DECAY SEPARATION...........................................................................................

PRED ACT PRED ACT PRED ACT PRED ACT _RED ACT.............................................................................................................

RANGE TIME--SEC 9224.92 9207.52 9207.42 921D.02 _548.G4 955D.SB 95_8.79 9550.83 149]_,87 14195.?2.............................................................................................................

5-IVB STAGE DRY 11596. i1563. 11596. 11563. I159B° 11583° 1159G. 11563. 11596. 115_3.

LCX IN TANK 60194, 60375. 60_30, 60211° 2068. 2244, '2037. 2212, 2037° 2212.

LOX 3ELOW TANK 166, 166. 18D. 18D. 180. 1GO. 180. 1_0° 16G. 16&-

LOX ULLAGE GAS 159o 156. 162. 158. 252. 269. 252. 269. 252. 269.

LH2 IN TANK 13151. 1319_. 13086. 1_11G° 892. 973. 979, 961, 879. B_l.

LH2 BELCW TANK 26. Z6, 26, 26, 26. 26. 26. 26. 22. 22°

LH2 ULLAGE GAS 15G. 143. 156. 14_. 272° 274° 272° 274. 272. 2 4.

ULLAEE MOTOR PROP C. O. O. 0. 0° O. 0. O. O. O.

APS PROPELLANT 183° 246° 183. 2WG° 179. 241, 179. 241. 144. 20_°

r,_ HELIUM IN BOTTLES 149° 163. 149. 163. 90, 99° 90° 99, 90. 99.OI START TANK GAS 2. 2. O. 0. 3. 3. 3. 3. 3. 3.

FROST C. 45. O, 45° D. 45. 0, 45. D. 45.

OTHER 25. 25. 25. 25. 25. 25- 25. 2S. 25, _5.

TOTAL _-IVB STAGE 858D9. 86D92o 85594. 85878, 15584. 15943° 15539° 15_99. 15486. 158_°

TOTAL INSTRU UNIT 1930. 1935° 193_. 1935o 1930° 19_5° 1930, 1935o 193_. 1935o

TOTAL SPACECRAFT 44679. 44572. 44679. 44572. 44579. 44572. 44679. 44572. G26, _24°.......................................... w .............................. . ...................................

TOTAL U_ER STAGE 46608° 46507. 4S&O_. 46507. 46508° 465D7. 46608. 46537, 2555, 25_I.

TOTAL VEHICLE 132417. 13ZGgO° 132_G3o 132385. &2192° _2450, 62147. 62407. 18341° 1B4OB.



T able 20 -8. T o ta l V eh icle M ass - S -IV B S econd B urn P hase - P ounds M ass

S-IVB S-IVB S-IVB S°IV_ SPACECRAFT

EVENTS IGNITION MAINSTAGE ENGINE CUTOFF END DECAY SEPARATION

PRED ACT PRED ACT PRED ACT PRED ACT PRED ACT

RANGE TIME--SEC 92D4,92 92D7,52 9207°42 9210,02 9548°64 9550,58 9548,79 955D,83 149DW,87 14185°72

S-IVB STAGE DRY 25564. 25492, 25564, 25492. _5564. 25492, 25564° 25492° 25564, 25492.

LOX IN TANK 132705, 133104, 132343, 132742, W5&O, 4947, q1491, 4877. 4491, 4877.

LOX BELOW TANK 3£7, 3BT, _9_. 397. 397. 397. 397° 397. 367. 367,

LOX ULLAGE 3AS 351, 343, 358, 349, 555, 593° 555, 593, SOS, 593.

LH2 IN TANK 28993° 29058, 2685D° 2e915, 1967, 2146, 1937, 7119, 1937, 2119,

LH2 _ELOW TANK 58, 58, 58, 58, 58° 55, 55, 58, 4B, 4B.

LH2 ULLAGE GAS 344. 316, 345. 317, 599, 603, 599° 604. 599, £DQ,

ULLAGE MOTOP PROP q, Oo O, Oo O, O° O. D, D° 3,

AP$ PROPELLANT 403. 5_2, _03, 542° 395. 531, 395° 531° 31B° 45_°

_ELIUM IN 50TTLES 329, 350, 329, 359, 198, 218, 19B. 218, 19S. 21B,

0 START TANK GAS 5. 5° I, I° ?. 7, 7. ? 7. 7,I °..-.' FFOST O, IO0, O, 100, O, 100. O. log, O. IOD°

0 OTMER 56, 56. 56. 56, 56. 56. 56. 5G. 56. 56.

TOTAL S-IV8 STAGE 189176, 189801. 188703. 189328. _4356. 35148° 34257. 35052. 3414_. 349_5.

TOTAL INSTRU UNIT 4254. _2&7. 4254. 4267. 4254° 4267. 4254. 4267. 425_. _267°

TOTAL SPACECRAFT 98580. 96264. 9850D. 95264. 785SD. 98264° 985D0. 9826q. 1350. 1380.

TOTAL UPPEP STAGE I_2754, It2531, IO2?SW, i02531, ICZTS_. I_2531° IC2754, I_2531, SE3_. 564T.

TOTAL VEHICLE 291930, 292332. 291_57. 291859, 137118. 137679. 137011. 137583, 39774, 405BZ,



Table 20-9. Flight Sequence Mass Summary

PREDZCT_O ACTUALMASS HISTORY K.G LBM KG LBH

S-IC STAGEe TOTAL 228203q. S03102q. 22795_,9. $0255qG.

S-IC/S-I1 INTERST J,GE eTOTAL $2G2o 11609. 5255. 11566.

S-II STAGEP TOTAL q8459D. IO68337. 48qlSG. 1067389,,,S-II/SIIVB INTERS TAGE 3EE5. 8061,. 36419 o 60qG.

S-IVa STAGEr TOTAl_ 118858° 252037. 119223. 2628_1.

INSTRUMENT UNIT 1930. 4254° 1935. 'qEG7oSPACECRAFT INCLUD:[NG LEG qBT31. 107q33. 48625° IOT2'OOo

JST FLT STG AT _GN 2945069. 6492766. 2942396° 6486873.

S-I(: THRUST BUILOUP -38916o -85795° -42013. -gZGZ2.

1ST FLT STG HOLDDWN ARM REL 2906153. 5qOG970° 2900383. G39.q251.

S-IC FROST -295o -GSO., -Z95° -6SO.

S-IC MAINSTAGE PROPELLANT -2069956. -_563q70. -206GGGG. -q554962.S-I(; N2 PURGE -17. -37. -17. -37°

S-IC INBD ENGINE ToO. PROP -O2q. -181G. -875o -1928.S-IC INBD ENG EXPENOEO PROP -185. -1t08o -[OO° -418.

S-II INSULATION PURGE GAS -S_.° -120,, -Sq. -]ZO°

S-II FROST -2D_. -450° -20q° -qso.

S-IVB FROST -136. -300o -91. -2100 °

IST FLT STAGE AT S-IC OECOS 83_483. 1839720° GIZSG2, 183Sq.IS6,

S-IC OTBD ENGINE T.O° PROP -3295° -7263° -3499° -7713oS-IC/S-IZ ULLAGE" RKT PROP -33. -73° -33* -73°

1ST FLT STAGE AT SICISII SEP 831155. 1832384° 82903h 1827700.

S-IC STAGE AT SEPARATION -Jl.GBsqG. -371584, -IGGSS6. -3GTZIG.

S-ICIS-II ]NTERS;TAGE SMALL -612. -]1SU. -611. -13q8.S-ICIS-IZ ULLAGE RKT PROP -83. -lOq. -83° -184.

2NO FLT STAGE AT S-II SSC GE,19]e2. lqS92GSo GGITS9. 1q58952°

S-II FUEL LEAD Oo rt° O. O.S-IC/S-II ULLAGE: RKT PROP -188. -414o -184o -ROE°

2NO FLT STAGE AT S-ZI IGN GG1T2q-. lq58851° 6615BS° 1q58546.

S-II T.B. PROPELLANT -593° -1306. -574. -IZGGoS-II START TANK -11. -25. -11° -26°

S-[C/S-II ULLAGE RKT PROP -313. -659° -]09. -662.

2NO FLT STAGE AT HAINSTAGE 660807. IqSS831. 6GOSOO. 1q56573°

S-II HAINSTAGE _, VENTING -4388nqo -987398,, -q38112. -9G5871°LAUNCH ESCAPE SYSTEM -4052. -8933. -q053. -8936.

S-I_/S-It INTERC.;TAGE LARGE -4032° -8890. -qrt]3° -B8_2.

S-II T.D. PROPELLANT -57. -127. -GO. -132°

2ND FLT STAGE AT S-I][ C.O.S. 213862. qT|qO5° 21Wq]2. q72741.S-II T.D° PROPELLANT -1G1. -355. -171° -378°

S-IVB ULLAGE PROPELLANI -2° -5. -2° -S.

END FLT GIG AT $)[IISIVB SEP 213699. 471125. 2"14258° _.7235B.S-ZI STAGE AT SEPARATION -_,4705. -9856G. -_,qOT2° -9914G°

S-IIIS-IVB INTERSTAGE-DRY -318S. -7021. -31G7° -6902.

S-II/S-IVB IS PROP -qBI° -lOGO. -q82. -1083o

S-IVB AFT FRAME -22. -qs. -22° -qS.

S-IVB ULLAGE PROPELLANT -h -3. -1. -3.°S-IVB 9ET PACKAGE -1. -3. -1° -3°

3RD FLT ST6 AT 1ST SSC 1SS30_. 3Gq432. 165E12. 385113°S-IVB ULLAGE PROPELLANT -4D. -88° -40. -8B°

S-IVG FUEL LEAD LOSS -O. -O. O° O°

2 0 - I I



Table 20-9. Flight Sequence Mass Summary(Continued)

P_EDICTED ACTUAL

MASS HISTORY KO LBM KG LBM

3RD FLT ST6 iT 1ST SIVS IGN lS5263, 36q3W3. 1655T3, 366925.S-IVB ULLAGE PROPELLANT -10. -22, -10. -22.

S-IVB START TANK -2. -q. -2. -q.

S-IVB T.B, PROPELLANT -198o -q3T. _205, -453*

3RO FLT STG AT MAINSTAGE IGSG5_. 363681° |_5365. 3&_SqG,S-IVB ULLAGE ROCKET CASES -6|. -135. -82° -13T.

S-IVB MAINSTAGE PROP -31186, -68753. -31q$9. -69356.

S-IVB APS PROPELLANT -h -2, -5. -10.

3RO FLT STG AT 1$T SIV_ COS 133806° 294991o 153830° 29504_.

S-IVB T°D. PROPELLANT -45. -99, -;q. -97.

3RD FLT ST6 AT END 1ST TD 133760. 294891. 133186. 29_94T,

S-IVB END PROP EXPENDED -28. -qD. -18. -qO.

S-IVB FUEL TANK LOSS -1188o -2619. -IOlq° -2Z36,S-IVG LOT TANK LOSS -20. -43. -83. -184,

S-IVB APS PROPELLANT -102o -225. -53. -116,S-IVB START TANK -1. -2. -1. -2.S-ZVB 02/H2 BURNER -7, -16, -7. -|6,

3RD FLT ST6 AT 2NO SSC 132_24° 2919_6° 132609. 292353.

S-IV8 FUEL LEAD LOSS -7° -16. -ID° -2|.

3RD FLT STG AT 2NO SIV8 ION 132417, 291930. 132600. 292332.S-IVB START TANK -2° -_. -2. -q,

S-IVB T°Bo PROPELLANT -213, -qG8. -213. -469.

3RD FLT STG AT MAINSTAGE 132293. 291_57. 132385. 291859.

S-IVB MAINSTAGE PROP -7OOOTo -154339° -69930. -|5q169.S-IVB APS PROPELLANT -;, -8. -6. -11,

3RD FLT STG AT 2NO SIVB COS 62192° 137110. 62qso. 137679.

S-IVB T.D° PROPELLANT -_S. -99. -_. -96.

3RO FLT ST6 AT END 2ND TO 62147. 137011. 62_0T, 137583.

S-IVB ENG PROP EXPENDED -18. -qD, -18. -_0.

S-IVB APS PROPELLANT -35. -Tie -35. -TT*

SPACECRAFT SEPARATED -q_OS3. -97120. -_39_&. -9688_.SPACECRAFT NOT SEPARATED -626° -1380. -626. -|380.

INSTRUMENT UNIT -1930. -425_. -1935. -4267.

S-IV8 STAGE AT SEPARATION -16_86. -3ql;O. -158_6. 34936.

20-12



Table 20-10. Mass C haracteristics Comparison

........................................................... _.................................................

WAGS L_?_51TUDINhL RADIAL BOLL MOMENT PITCH MONFNI TAW MOMENT

¢._, IX _I_,l C.G, UF INERTIA OF INERTIA OF INERTIa

1;V[ NT ........... =..........................................................................

KILO 0/0 NLIL_S MFTENS KG-142 010 KG-M2 010 KG-H2 010POUNDS OEV. I NCHF S 01;L TA INCHFS OILTA _IO-G DI¥• X|D-G O_Vo X|D-O OGVo

................................................ - ..........................................................|IIqqT. q, qOl .07GI

PRFO 79q200, ITO. I 2. T659 2.GW2 IT.l R5 IT.|OO............................................

9-TC %T&OE _qY 1333qS. q.qoI .OOO °oT03 .O(]O0

ACTUAL 793979, -*07 3TO,1 *GO 2, TG59 *0_0_ 2,GOD -tOT 17,172 -.O? l?.On? -,OT...........................................................................................................

52_2 • k 1.G2 3 • 19 _lEPRF(_ 1 IG00, _638. T 6,0877 ,1]_ *OG| .DO|

S-TO/S-I I INT1; g .............................................

STAGE* TOTAL 5295. _I,EZO ,U03 *|5G3 .0017

_CTUAL 11585. -.12 IG38.8 o10 6,1555 oOGTR *IIW -.l_ ,080 -ol2 .001 -.12.......................................................................................................

3B269. _q.ll $ •ll IG

pR ED OqIGT. I 89q. 2 q, 39 32 .G 33 2.1 TE 2 .| 80.....................................

9-TI CTAOF ._RY 38226, Itq_090 -,023 .1116 ,(_00

ACIUAL 8_273, -,1] 1893._ -,9C _o3932 •COOl] ,G32 -.|1 2.1TO -,|1 2o185 -*11........................................................................................................

3GGG. EG.BGO ._573PR fO 80 81 • 2 G9 2.9 2. 29 61 .0 G5 *0 _I .0 _el

S-TI/rcIV_ INTER ...............................................STAGE_ TOTAL $650* I;5.938 •DIG . "_998 .0025

ACTUAL BOOSt -*_q 2995,9 3,0(2 2",_537 •09TO ,GO5 -,_0 ,ON3 -°qq ,Oq_ -,N_

llG81. T2•SOl •1978

PR ED 2 GT $0 • 2 OG8.3 7 • 7878 °0 _2 °305 .308............................................

%-TVO ST AGF,ORY I IGla9, T2, GOI ,0110 *19TB ,OOOOACTUAL 25GOO. -°_? _58.3 °DO 7_TOT8 °OGDO .COZ -*2T .90_ -t_T ,30G --_27

1930° O2*_tI9 • 3511

PREO q25_, 32_q,T 1 ]°R292 -019 .OIO .009VFHICLG INGT_ UM1; NT .............................................

UNIT 193G, B2. q15 .DUO ,3511 ,GOOD

ACTUAL _2ST. .31 3Zqq,T .OOl 3_ 82 32 ,0000 *Ol9 ,31 ,DID ,31 ,009 ,31

98731, 91,G5I ,1085

PROD 10T_33, 3GO8. q q. 2320 ,09G I,G52 I,GG5

5P ACACIA F"TTOT AL q BE _'S. S1,GG8 .COS ,1099 ,GOlq

ACTUAL lOT2OO* -,2] I_O8,G °_O _._2G? .OGqT ,088 -1•70 l.Gq_ -,Zl IoSSO -,30..........................................................................................................

_.'9_5070 ° 30.3L 1 • _O _2PNEO G_92TGE. llgI. q ,IGqO 3o78G 9lG,?q_ _15 ,G59

IKT FLIGHT STAGE ......................................

AT IGNITION 29q 2397. _0.317 .DOG .nOq2 •OOO0

ACTUAL GqOG870, -*US II53•G °29 ,liNG _O_QO 3,719 -,IG 915,GG_ -•00 _1Go592 -,O0........................................................ . ...............................................

29061530 30.25T •00q2PRGO GqO6970, I191.2 *IEqO 3._B9 91_.G lq _ lq *_2_

|_;T FLIGHT _TAGF .......................................

AT HOLODOWN _RH 290038tt, 30,259 .002 *l_t_2 •0000RTLFA_E ACTUAL 639_252. -°19 IISh_ .08 •]6(;0 ,00_ 3.782 -*lG 91_,23G -,03 5Jq°_S -_03......................................................................................................

839_89 • 9_, 2| G ° '31 ql

PROD 1_39720, 1919,5 •SSq7 3°779 q_S,8IG q _5,T_9I_T FL][GHT STA_C ...........................................AT OUTROAOD FNGIN_ O325G- _. _#G°932 .1]7 •r)]_#2 ,0001

CUTOFE SIGNAL ACTUAL 1835_8G, -,2_ 1829,1 _°G9 ,GGOq *OOSE 3.T67 -.15 qql.9_O -_OG q_1,853 --,OG.....................................................................................................

831196, _16,359 •01 ql

PREO ]83238tt ° 1025.2 _GGqT 3°TTI _ql•_ 991,3_S}_T FLIGHT _CTA_'C ............................................

AT 5EP EOAT ION R29031 ° _O._81 .122 • f3I t12 •0001ACTUAL 1827700, -,2!_' |_30eO q°79 ,GG'iT'_ ,005£ 3.TGS -o1_ _TT._8$ -,89 qST.Iq8 -.89

.... .... ... .... .... .... .... .... .... .... ... .... .... .... .... .. -_ - .... .... .... .... .... .... .... ... .... .... ...

EEISI2. 55.750 ,olq2P_EO |_592G5• 2IS_,q • _G (]l I ,021_ I _,893 |lq.OT9

2NO FLIGHT _Tk_E ......................................AT GIANT _EOU_NC£ GGITTO, GG.?GZ .012 •_Iq2 -•0000

COM_4ANO ACTUAL lOGO952, -•02 _1_5,_ ._8 ._600 -*C_01 I°OIR -,2_ ]3q.8qB °OZ 13_,G61 ,02......................................................... _-_-_.. .........................................

GGOEOB ° 55°7E_ *01 tt2

_RGO lqGGO_l , _lSGoq ._GOI l,OO9 | 3_°T ltt I ]q oT202NO FLIGHT STAGF .............................................

AT MATN!; TAOr _6UGgl * _5.773 .011 . (_[ it2 -, r)o O_ACTUAL IeGGGT], -,01 _7199*8 .qct •SGCO -,1_0| I.GOG -*29 I]q*?_ °02 |_q,T'q_ ,O_

.........................................................................................................

_13A62* _0,792 •_ |S

PREO eTl_GG, 2787,1 1,£325 ,905 _9 uS08 _5.51_

.2N0 FLIGHT 5TAGF ...............................................•AT CUTOFF sIGNAL ?It_432. 70,75_ -*03B ,q_IS ,[_00

ACTUAL qT27_l. .27 7785,_ -1,G0 I,F325 _000{1 .gO2 -,2T _9o69_ ,ql _S°YO6 •tl3

20-I 3



T able 20 -1 0 . M ass Characteristics Comparison (Continued)

.............................................................................................................

MArc<; LI)_(;ITUDINAL RAOIAL ROLL MOMENT PITCH MOHENT VAW MOMENTC'.f.. |)r STA.) C,G* OF INERTIA OF INTRTIA OF INERTIA

t VE'qT ...................................................................................

K]L G 0/0 _t_T[R S MFTERS KG-MZ 0/0 KG-M2 0/0 KG-M? 0/0POUKIOS DEV. ] _dCHEG OELTA IN CH I'-G OELTA XIO-G nEV. X]D-G OIV. X|O-G GEVo

Yl"_Gqo. TO.B| | .nt_ 17

P_EO q71125, _ 787. B ]o _q 25 .905 Q5 .q 07 415°COT_Nr3 FLIGHT '_TA_E ........................... __ ...............AT SEPARATION 71¢,258. TO. 77=_ -.03G ° tt_t15 -. 000_

ACTUAL _72358, *27 276G.q -l°kI l*GI?5 -.01flO ,9.O2 -.27 qG.SGO *uO q5.5_3 ._|............................................................................................................

1_fi30_. 77.08_ .P377

r'qFO 36_2, 303_.8 1.qGqI .195 13.31t2 13,339lqO FLIGHT STA_E ........................... __ ...............

AT [ST START SAG- LGGG]I, 77.QG3 -.020 ire37G .f)001

Ur_cE COM_tANt3 ACTUAL _GG113* ,]9 303_.0 -o81 1,qE72 .0031 ,lGq -._G tToI31 -.08 13o328 -°08.........................................................................................................

Lf_S2Gt=• 77.081 °037T

PRED 1GctIqI. 303_°7 1. q8 tel .195 13.3q3 13.3lEO_0 FLIGHT _TAGE ...........................................

AT LST IO_'/ITION 165573. 77.O_fI -*0;_O .nIT8 .O(_G!

ACTUAL 365(37G. ,1_ 3033._ -*G1 1*16872 °O03t .19q -*tAG 1T,31I -_0_ z 13o328 -.68.............................................. _ .....................................................

IGSOGq. IT.DOlt • 0377

Pf_EG IG 3_81 * 363q. 8 1, q8 ql .19_ 13.3 ql l1.337

31_ FLIGHT _TAGE ....................................AT ]_I MAIN_TAfiE 1E535G. 7T.OG_ -.U21 °DI_qO °0003

ACTUAL 3GqGqG, ,]G 3OIq°O -°83 l,qq57 *DIIG ,|_;t -°qG 1"{.3_1 -.GO 13.32G -*08

13300G° T_°015 . OqG2t

P_ED 79q991 • 3071°5 1,82 _tt *|Gq 1_.518 1_,51_XPO FLIGHT STAG[ ............................................

_T I_T CUTOFF GIG- 133830. 77._9_ -*O21 ,(IqGG .OOOT

NAL ACTUAL 29GOutt. .O_ 3070.G -.83 1,GI3G .0115 o|_3 -°GO 12.512 -.0_ l_*GO9 -°Oq...................................................................................................

133761. 7_.017 .OqG3

PLIED _9q8_1. 3071.5 1. 822tt ._Oq l_.GIT 12.G1 tlIP9 FLIGHT _TAG[ ................................

AT |ST EN_ THRUST 1337qG, 77,99_ -,021 ,OqGG *0063

O_CAY_ _TART COAGTACTUAL 29q9_7, °02 ]070°7 -*83 1._338 ,OILS ,1_3 -,50 12,511 -*Oq 12o508 -°0_.................................................................................................

13Zq25. 78.0_7 ,OqGq

Pf_EO 2_lGqG. 3671.9 1,8250 .193 12.5 1_ lZ,GlO]P_ FLIGHT STAGE .........................................AT 2NO START GAD- 132G10° 78.OOG -.OZI .OqG7 ._6oq

uFNCE cOMMAND ACTUAL 292353. *Iq 3071,1 -,83 1,8392 ,Olq? .192 -.11 12,50_ --Oq 12,50q -*OG.......................................................................................................

13Z_18. 78.6ZG .04 6q

PREO 291930. 30TI.9 1.8250 .193 L_.5 Iq 12.512_0 FLIGHT _TAOE ..........................................

AT ZNO IGNITION 112G60. ?O. OOG -°026 .O_G7 o(]00_

ACTUAL 29__33Z° ,1_ 3071,1 -.81 1.83_2 *01q2 *1_2 --*11 IZ°509 -,0_ 12.50G -,05.................................................................................................

13_203. 78.031 , (1_Gq

PREO _91qG?, 3672, Z 1,8250 ,193 12 .$ [_ 1_ .SOT_PO FLIGHT STAGE .........................................

AT 2N0 MAINSTA_E 13238G* 78.012 -.OZ] .O_tGT .060=I

ACTUAL ZG1859. ,lq 3071.3 -.83 1*8392 .Olq? .192 -*ll 12*G63 -*Oq 1Z°501 -°O5..............................................................................................

62192. 85.722 *OgTq

PREO 137LO9. 337q°g 3.8355 ,192 5,315 5,3123GO FLIGHT GTAG{ ............................................

AT 2N[ t CUTOFF G_qGo* 86.631 -.OG1 _09?6 .GO0|

_EGN_L ACTUAL 137678, ,q2 33T1°3 -3,59 3,'8_13 .OOq8 ,19_ -,l_ 5.391 [*_3 5,388 1°q$................................................... __._____ _ ..........................................

621q7. 95,73q *OOTq

PGEO 137010° 337_,3 3° 83_ ,1_2 $ ,I _Wt 5,301lq_ FLIGHT STAGE ....................... _ ..............

aT 2NO [N_ THRUST G2q07, AS°GEt2 -.092 °097G .OOOl

DECAY ACTUAL 1375G_, ,_2 3_71,7 -3°G1 3,8q13 ,noqR *192 -,lq 5.3GO I°q5 5.3IT ]°_t.q.......................................................................................................

3ZOIq. 7B*_77 .OTSG

mREO 70S 22 • 3089.6 3* O153 • ! 36 1.6 33 ! °6 _8

3Z353. 78.387 -*090 .OT7q .000_

ACTUAL T132G* l,OO 3086.1 -3.56 3. Oq38 *0305 .I 36 -.32 1.655 l. Ik l.G_G 1.31.......................................................... ___ ......................................

60906, 85.10_ • 1262

P_EO 13qZII, 335_,G q, OG_G °83 q,TSB q,75_C$_ DOCK19 ... .. ... ... ... .. ... ... ... .. ... ... ... .. ... .

6117G* 8G.073 -. 08.G .l_EI -. OOOl

ACTUAL 13Z1875° ,q5 33q9,$ -],q8 q, oEqO -,(]USE ,18Z -,2_ q,831 1,53 q,A2q 1*32.........................................................................................................

IGOkl. 73, G03 .1377

P_EO 39773. 2097, T 6. q_O5 .108 .620 *Gl7SPACEC#A F T _;EP ..............................................

A_*ATEO 19_U6, 73._6 -°Oq7 ,13ql -, O030

ACTUAL qOG81, 2,0_ 289_,9 -I,B5 S.$OqO -.116G ,108 -*q7 *G27 1,13 *G2q 1.06......................................................................................................

20-14



S ECTIO N 21

M IS SIO N O B JE CT IVE S A CCO M PLIS HM EN T

Table 21-I presentsthe MSFC AS-505major flightobjectivesand detailedtest objectives as defined in the Saturn V Mission Implementation Plan,

Mission F. An assessment of the degree of accomplishment of each objec-

tive is shown. Discussion supporting the assessment can be found in the

indicated sections of the Saturn V Launch Vehicle Flight Evaluation

R eport - AS-505, Apollo lO Mission.

Table 21-1. Mission Objectives Accomplishment Summary

MSFCMAJORFLIGHTOBJECTIVES(MFO) DEGREE PARAGRAPH

NO. AND MSFC SECONDARYDETAILED OF DISCREPANCIES IN WHICH

TESTOBJECTIVES(DTO)* ACCOMPLISHMENT DISCUSSED

1 Demonstratelaunchvehiclecapability Complete None 4.3to injectthe Spacecraft(SC)ontothe specifiedtranslunartrajectory,(MFO)

2 Demonstratelaunchvehiclecapability Complete None II.5.4

to maintaina specifiedattitudefor 12.6Transposition,Dockingand SC Eject-ion (TD&E)maneuver.(MFO)

3 DemonstrateS-IVBpropellantdump and Complete None 7.13safing.(MFO)

4 Yerify J-2 enginemodifications. Complete None 7.3(DO 93.3.3

5 ConfirmJ-2 engine environmentin Complete None 9.3,16.3.2

S-II and S-IVB stages.(DTO) 17.3

6 Confirmlaunchvehiclelongitudinal Complete None 9.2.3

oscillationenvironmentduringS-ICburn period.(DTO)

7 Verifythatmodificationsincorpora- Complete None 9.2.3

ted in the S-IC stage suppress low

frequencylongitudinaloscillations.(DrO)

8 Confirmlaunchvehiclelongitudinal Complete None 9.2.3oscillationenvironmentduringS -II stage burn period. (DTO )

9 DemonstratehatearlyS-Ifcenter Complete None 6.3

enginecutoffsuppresses-Ifstage 9.2,3low frequencylongitudinal oscilla-

t io ns . ( DT O )

*Therewere no MSFC principaltest objectives;a11 testobjectiveswere classified

as secondary.

21-I/21-2



S ECT IO N 22

FA IL UR ES , A NO M AL IE S A ND DE VIA TIO NS

22.1 S UM M A R Y

E valuation of the launch vehicle performance during the A S -505 flight

test revealed one area of concern with a mission criticality categoryof three. A ction is planned to prevent reoccurrence of this problem onfuture flights.

22.2 S Y ST E M F A IL UR E S A N D A N O MA L IE S

T able 22-I defines the criticality categories assigned to the failuresand anomalies "listed in Table 22-2, which complies with Apollo ProgramDirective N o. "19 . R eference paragraph numbers are given for sectionsin which the specific problem area is discussed in m ore deta il.

Table 22-I. Hardware Criticality Categories For Flight Hardware

CATEGORY DESCRIPTION

1 H ardw are fa ilu re w h ich resu lts in lo ss o f life o f a n ycrew member. This includes normally passive systemssuch as the E m ergency Detection S ystem (E DS ), L aunch

E scape S ystem (L E S ), etc.

2 H a rdwa re fa ilu re w h ich resu lts in a bo rt o f m iss io nbut does not cause loss of life.

3 H a rdw a re fa ilu re w h ich w ill n o t resu lt in a bo rt o fmission nor cause loss of life.

22.3 S YS TE M DE VIA T IO NS

N ine system deviations occurred without any significant effects on theflight or operation of that particu lar system. T able 22-3 presents thesedeviations with the recommended corrective actions and a reference to

the paragraphs contain ing further discussion of the deviation. T hesedevia tions are of no m ajor concern, but are presented in order to com plete

the summary of deviations experienced on A S -505.

2 2 - I



Table 22-2. Summary of Failures and Anomalies

FAILURE/AtlOF.IALYDERTIFICATtOR RECO[_IENOEDORRECTIVECTIO!_

MISSION TIME VEHICLE

VE!!ICLE EFFECTOrlCR[TI- EFFECTOIJ OCCURRENCE ACTION EEFEC- _ARAG_AP_

E ! ST EI! DESC R IPT ION(CAUSE) N ! SS ION CALITY R EXT IIIS SIO _ (_ IAN GEIME) DESC R IPT ION ST ATUS T IVITY R E FE R E !IC E

S-IVG #_xiIiaryhydraulic _o_e 3 TIo_e 9425 l_specti_of sp_i_g Closed AS-5O6 _.7:_ydraulicspumpStopoed_roducing seconds guidesFor proper for and

fullpressureduring filletradiusand AS-506 SubsSecondburn. (Susoect propereock_le_]fai_uTeOf a_xi_i_T_ $_a_ness._ns_ect

hydraulicO_D comoensatororinescocpensatorprinc tOinsureall

guide.) tolerancesremet.

Table 22-3. Summary of Deviations

CORRECTIVE ACTION PARAGRAPHVEHICLE DEVIATION PROBABLECAUSESYSTEM BEINGONSIDERED REFERENCE

S-IC Lo_¢erformancef Unknown None.Average 5.3

Propulsio_ engine No. l thrust thrust over fullreduced to standard burn was more

conditions was 97,000 nominal.

N e wt o ns (22 ,0 00 l bf )

bel ow p re di ct ed .

S-IC UnexplainedOXsuction Unknown None, Similaroccur- 5,6.2

Propulsion duct pressu_ decay of fences during AS-503

engine No. 5 after CECO. and AS-504 with noeffect on mission,

S-[[ Slightly sharper pres- Leak through J_2 None. Decay rate 6,2

Propulsion sure decay of engine engine helium _eturned to normal

No. 5 helium tank regulator at 60 seconds

pressure than expected, after ESC.after ESC,

S-IVB Astronauts reported Data indicate S-IVB had None anticipated, 9,2.3

Propulsion/ mild low frequency typical buildup and decay but MDAC ECP 3218

Mechanical oscillations (12 to ]9 periods of very mild 12 adds 5 measurements

hertz) dur}ng first to 19 hertz oscillations for stabillty modela_d secondb_rns, withoutindicationsof analysesaBdflight

propulsion/structural evaluation of low

coupling, frequency oscilla-tlons.

S-IVB Astronauts reported Cycling of the LH2 tank Test program at 9,2,3

Propulsion/ noisy low level vibra- NPV valves AEDC underway to

Mechanical tions during latter confirm interacting

part of second burn of LHB tank NPVwhich were superimposed valves.on the 12 to 19 hertz

vibrations.

S-IVB Unexpected increase in Unknown None. Has been ex- 8.6

Propulsion/ S-IVB engine driven oerlenced on other

Hydraulics hydraulic pump outlet systems (F-IO0 Air-

pressure (3 percent) craft) and is not

shortly after second considered a problem.b ur n s ta rt .

S-IVB APE Module No, l helium Unknown. Similar problem Being investigated. 7.12

Auxiliary/ supply pressure decay on AS-504 res_lted in Leakage rabe insuffi-

Propulsion at approximately 23,400 cbanqe of seal material cient to impact

System seconds, and additional leak check mission.at KS C.

IU/RF Erratic signal strength Malfunction of coaxial Mope, (Coaxial switch 1g,5.3.2

System at receiving station switch, to be replaced on

beginning at 23,601 AS-507 per previouslyseconds, plannedCP).Omni

dir ec ti on al a nt en na

s ys te m p ro vi de d s uf fi -

ci en t s ign al s tr en gt h

t o m ai nt ai n s at is fa c-

tory communications,

IU/GN2 Sharp drop in IU inlet Opening in the purqe Installation of dual 18.4

Purge pressure and increased duct, clamps on umbilical

System flowrate at -9.8 hours connection boot withaccompanied by a tom- increased clal_Dtoroue.

p le te ID es o f p re ss ur e

t o t he R ad io -l so to pe

Thermo-Electrical

_e ner at o_ on t he L M.

22-2







A PP E N DIX A

A T M O S P H E R E

A .I S U M M A R Y

This appendix presents a sunwnary of the atmospheric environment atlaunch tim e of the A S -50 5. T he fo rm a t of these da ta is sim ilar to tha t

presented on previous launches of S aturn vehicles to perm it com parisons.

S urface and upper winds, and therm odynam ic da ta near the launch tim e aregiven.

A .2 G E N E R A L A T M O S P H E R IC C O NDIT IO NS A T L A U N C H T IM E

A h igh pressure cell, in the A tlan tic O cean off the N ew E ng land coast,

caused southeasterly surfa ce w inds and brought m oisture in to the C apeKennedy, F lorida area, which contribu ted to the overcast conditions

during launch.

A .3 S U R F A C E O B S E R V A T IO NS A T L A U N C H T IM E

At launch time_, skies were overcast with 4/10 cumulus at 0.7 kilometer(2200 ft), 2/10 altostratus at 3.4 kilometers (II,000 ft) and I0/I0

cirrus at _n unknown altitude. Surface observations at launch time aresum m arized in T able A -I. S o la r radia tion da ta are g iven in T able A -2.

T able A -I. S urface O bservations at A S -50 5 L aunch T im e

SKYCOVER WINDTIME PRES- TEM- POINT VISI-

LOCATION AFTERSURE PERATUREDEW BILITY AMOUNT TYPE HEIGHT SPEEDT-O N/CM2 °K °K KM (TENTHS) OF BASE M/S DIR

(MIN) CPSIA) (°F) (°F) (STATMI) M (FT) (KNOTS) (OEG)

Kennedypace 0 I0.190 299.82 295.3_ 18 4 Cumulus 671 5.7 130

Center,Station (14.78) (80.0) (72.0 (II) (2200) (If.O)Merritt Island,Florida 2 Alto- E3350

cumulus EllO00)

l0 Cirres high

CapeKennedy I0 10.184 300.25 295.2! ........ 6.0 120

Rawinsonde (14.77) (80.8) (71.8 (11.7)Measurements

Pad39B 0 ............... 8.2 125

LightpoleE (16.0)(20.I m)*

* Abo ve N at ur al G ra de

A - l



T able A -2. S olar R adiation at A S -50 5 L aunch T ime, L aunch P ad 39B

HOURNDING TOTAL NORMAL DIFFUSE

EST HORIZONTAL INCIDENT SKY 2DATE G-CAL/CM G-CAL/CM G-CAL/CM(MIN) (MIN) (MIN)

May17, 1969 0600 0.01 0.01 0.010700 0.16 0.33 0.06

0800 0.41 0.53 0.14

0900 0.73 0.70 0.25

I000 1.04 0.81 0.36

II00 1.13 0.46 0.70

1200 1.19 0.33 0.87

1300 1.42 0.57 0.87

1400 1.34 0.50 0.89

1500 1.20 0.41 0.881600 0.96 0.28 0.78

1700 0.64 0.19 0.56

1800 0.33 0.II 0.31

1900 0.07 0.03 0.07

May18, 1969 0600 0.01 0.00 0.010700 0.14 0.16 0.09

0800 0.41 0.46 0.17

0900 0.74 0.61 0.32

I000 1.04 0.68 0.47

II00 1.19 0.58 0.651200 1.15 0.26 0.89

1300 1.36 0.37 1.00

1400 0.94 0.09 0.86

1500 0.54 0.02 0.52

A .4 UP P E R A IR M EA S U R E M EN T S

Data were used from four of the upper a ir w ind system s to com pile the

fina l m eteoro log ica l tape. T able A -3 sum m arizes the da ta system s used.It was necessary to use interpolated wind and thermodynamic data from

7 to 7 0 kilom eters (1 87 ,0 0 0 to 229 ,6 6 0 ft).

A.4.1 WindSpeed

Wind speed increased with altitude, reaching a speed of 42.5 m/s (82.6knots) at 1 4.1 8 k ilom eters (46 ,520 ft). W ind speeds at h igher a ltitudes

A - 2



T a ble A -3 . S ystem s U sed to M easure U pper A ir W in d D ata fo r A S -50 5

R EL EA SE T IM EP O RT IO N O F D A T A US E D

TYPEOFDATA TIME START ENDT IM E A F T E R

(UT) T-O ALTITUDE TIME TIME

(MIN) M AFTERALTITUDE AFTERT-O M T-O

(FT) (MIN) (FT) (MIN)

FPS-16Jimsphere 1704 15 0 15 15,750 69

( 5 1 , 6 7 0 )

Rawinsonde 1659 I0 16,000 62 24,750 91(52,490) (81,200)

Loki Dart 1928 159 56,750 159 25,000 187

(186,190) (82,020)

Viper Dart 2030 221 89,750 221 70,250 222

(294,450) (230,480)

w ere less th an th is peak, excep t near 9 0 k ilom eters (29 5,27 0 ft) a lti-

tude. S ee F igure A -I fo r m ore in fo rm ation o f the w ind speeds.

A .4.2 W ind Direction

The surface wind was from tile southeast, but shifted through the southto w esterly a t 1 4.0 k ilo m eters (45,9 30 ft) a ltitu de. A bo ve th is a lti-

tude winds sh ifted th rough the north and stayed generally from the east

above 18.0 kilometers (59,050 ft), as shown in Figure A-2.

A .4.3 P itch W ind C o m pon en t

The surface pitch wind speed component was a head wind of 4.0 m/s (7.8kno ts) and sh ifted to a ta il w ind by 3.0 kilom eters (9 840 ft) a ltitude.A maximum tail wind of 40.8 m/s (79.3 knots) was observed at 13.8 kilo-meters (45,280 ft) altitude. Head winds were observed from 16.9 kilo-m eters (55 ,450 ft) to 83.5 k ilom eters (27 3 ,9 50 ft) a ltitude , w ith a

peak head wind of 39.5 m/s (76.8 knots) at 71.0 kilometers (232,940 ft)

a ltitu de. S ee F igu re A -3.

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Table A-4. Maximum Wind Speed in High Dynamic Pressure Region for

Apollo/Saturn 501 through Apollo/Saturn 505 Vehicles

MAXIMUMWIND MAXIMUMWINDCOMPONENTS

V E H I C L ENUMBER SPEED ALT PITCH(Wx) ALT YAW(Wz) ALT

M/S DIR KM M/S KM M/S KM

(KNOTS) (DEG) (FT) (KNOTS) (FT) (KNOTS) (FT)

AS-501 26.0 273 ll.50 24.3 ll.5C 12.9 9.00

(50.5) (37,700) (47.2) (37,700) (25.I)(29,500)

AS-502 27.1 255 12.00 27.| 12.0C 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 15.1C 22.6 15.80

(67.6) (49,900) (60.6) (49,500) (43.9)(51,800)

AS-504 76.2 264 II.73 74.5 ll.7C 21.7 II.43(148.1) (38,480) (144.8) (38,390) (42.2)(37,500)

AS-505 42.5 270 14.18 40.8 13.8Z 18.7 14.85

(82.6) (46,520) (79.3) (45,280) (36.3)(48,720)

Table A-5. Extreme Wind Shear Values in the High Dynamic Pressure Region

for Apollo/Saturn 501 through Apollo/Saturn 505 Vehicles

(Ah = I000 m)

PITCHPLANE YAWPLANE

VEHICLE ALTITUDE ALTITUDE

NUMBER SHEAR KM SHEA_ KM

(SEC-I) (FT) (SEC-I) (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(52,500) (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

(50,200) (50,950)

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A.5.3 Atmospheric Density

Atmospheric density deviations were small, being less than 4 percent de-

viation from the PRA-63 to 36 kilometers (118,110 ft) altitude. Since 1

density gen era lly fo llows pressure p atterns, there was an increase in

ensity differences above 36 kilom eters (1 1 8,1 1 0 ft) a ltitude, with a

peak percentage difference a t 27 .6 percent from the P R A -6 3 a t 80 .5 k ilo-meters(264,100t) altitude. I

A .5.4 O ptica l Index o f R efraction

A t the su rface , the O p tica l Index of R efraction w as 9 .81 (n - l) x 1 0 -6

un its lower than the corresponding va lue of the P R A -6 3. T he devia tion

decreased w ith a ltitude, becom ing a m a xim um of 1 .9 2 (n -l) x 1 0 -6 grea te rthan the corresponding va lue o f the P R A -6 3 a t 1 3.3 k ilom eters (43,6 30 ft).

A bove th is a ltitude th e O ptica l Index of R efractio n approxim ates the

P RA -63 values.

A .6 C O M P A R IS O N O F S E L E C T E D A T M O S P H E R IC DA T A F O R S A T U R N V L A U N CH E S

A sum m ary o f the a tm o sph eric da ta fo r each S a tu rn V la un ch is sh ow n inT able A -6 .

Table A-6. Selected Atmospheric Observations for Apollo/Saturn 501 throughApollo/Saturn 505 Vehicle Launches at Kennedy Space Center, Florida

VEHICLE DATA SURFACE DATA INFLIGHT CONOIIIONS

VEHICLE DATE TII_ LAUNCH PRESSURE IEMPERA- RELAIIVE WIND" CLOUDS MAXIMUMW]ND IN 8-16 KH LAYER

_UMBER NEAREST COMPLEX N/CIA2_ TE°C HUMIDITY SPEED D[2EDTIOH ALTITUDE SPEED DIRECTIONHINUTE PERCENI M/S DEG M H/S DED

AS_SDI 9 Nov 67 I)700EST!¸ 32A I0.26] I/.6 55 8.0 _0 I/]D ¢umu]u$ II.5D 26.0 2/3

AS-502 4 Apr 68 ONO0 EST 32A 10.200 20.9 83 5.4 I32 5/10 stratocumulus 13.00 27.I 255

AS-5D3 21 Dec 60 ()75|EST 39A ]0.20/ I5.0 88 1.0 360 4/I0 cirru_ ]5.22 S4.B 284

AS-SD4 3 Mar 69 i|OO EST 39A 10.995 19.6 6] 6.9 160 I0/I0 sCrato- II.73 76.2 264¢_ulus

AS-S05 ID Nay 69 1149 EDT 390 I0.190 26.7 15 8.2 ]25 4/IO cumulus,2/I0 14._B 42.5 2/0altocumulus, 10/]0

cirrus

*InstantaneousreadinDs from charts at T-D from anemometerson launch pad at I8.3m (60.0 ft)on launch cc_p|ex32 (A&B). Heights of aneometer_

are above naturdl grade.

A- 13/A- 14



A PP EN DIX B

A S-50 5 S IGN IF ICA NT CO NFIGUR ATIO N CHA NGE S

B .I IN TR O DUCTIO N

A S -50 5, fifth flight of the S aturn V series, was the third m anned A polloS aturn V vehicle. T he A S -50 5 launch vehicle was configured the sam e asA S -50 4 with significant exceptions as shown in T ables B -I through B -4.

The A S -50 5 A pollo I0 spacecraft structure and components were essentia llyunchanged from the A S -50 4 A pollo 9 configurations. T he basic A S -50 4vehicle descrip tion is presented in A ppendix B of the S aturn V L aunchVehicle F light E valuation R eport A S -504, A pollo 9 M ission, M P R -S A T-FE -69 -4.

B - I



Table B-I. S-IC Significant Configuration Changes

SYSTEM CHANGE REASON

S tructures Integral machine fittings Increase reliability ofreplace welded fittings on bulkheads.

LOX and fuel tank bulkheads.

Data DeletedODOP transponderand ODOPsystem no longer

instrumentation, requiredfortracking.

Ten acoustic measurements To determine effect of

added to intertank, exhaust plume on vehicle.

Deleted fuel tank slosh R &D instrumentation which

probes, isno longerrequired.

Table B-2. S-II Significant Configuration Changes


S tructures Incorporate redesign of Improve weldability andLH2 feedlineelbows, reliability.

Propellant Use PU system open loop Improve reliability.Management mode (was closed loop

mode on S -II-4) .

Propulsion Command early cutoff of Avoid low frequency

center engine (No. 5) by oscillations e xperienced

switchselector, duringflightsof S-II-3and S-II-4.

Launch Add redundantventsystem Assureventingof vent

Vehicle to $7-41 for ventvalve valveactuationpressure

Ground actuationsystem, prior to -15 secondsto

Support avoidinadvertentpeningEquipment of stage vent valvesand

consequent loss of ullage

pressures.

B-2



T able B -3. S -IVB S ign ificant Configura tion Changes


Instrum enta tion A S -50 2 a nom alies instru - P rogram requires A S -50 2m enta tion package is anom alies instrum en tationinstalled and incorporates on AS-503 and AS-505the FM/FMand single side- stages only.band telemetry systems andadditional measurements.

Propulsion Two S-IVB J-2 engine burns. Normal TLI mission.F irst restart propellantta nk re pr es su riza tio n

performed by 02/H2 burnerwith ambient spheres asbackup.

R e m ov e an d inspe ct the T o e lim in a te lea ka g ebu lkh ea d fitt in g s a n d o f A P S h e liu m w h ich w a stube asse m bly fla re s in observ ed during the

the APS high pressure AS-504 flight.system, and the temperaturetransducer fittings. Re-install the fittings usingM S -28778, N itrile R ubber,90 durometer hardness "0 "

r i n g s .

Dele te on e L H2 tank am bient P aylo ad sav ings.repressurization bottle.

C o n n ect th e 2 L O X ta n k In crea se th e reservesam bient rep ressuriza tion capability for propu lsionbottles to the stage dumping and stage safing.pneumatics bottle.

Thermo- Thenumber of cold plates To accommodatethe

condition ing loca ted in the S -IV B fo rward additiona l electrica l

System skirt increased from 5 to 8. components.

B -3



Table B-4. IU Significant Configuration Changes


Electrical The S-]VOEngineCutoffEnableCircuitry The S-IVB restartrequirementafterwas not installedon S-IU-505. Spacecraftseparationisnot required

for S-IU-SOS.

Penllanentix to isolateFlightControl CCSgeneratednoiseon 6D41 bus wasComputer (FCC) from CCSgenerated noise fed back through ESEpower buses to theduring ground checkout. FCC(60311. Procedural change had been

usedto operateESE + bD211(CCS) topreventnoise fromCCS feedingbackto the FCC.

CableModifications Minimumcableand networkmodifications

were made to facilltatedisablingUHFcont.01circuitryand rerouteIU and

S-IVBpCNslgnals.

Enviror_entel PreflightONp/AirPurgeDuct modifications Additionalductswere routedto theRTGControl at Locations-In and 23, Fuel Cask located in the LM Descent

Stageto providepreflight cooling.

The FCCM/M supply_as disconnected. The possibilityof M/W leakagewaseliminated;FCC coolingnot requlred.

Guidance LVDAP-23 Circuitchanges. LVDA circuitchangesweremade toinhibit recurren't generationof ErrorTimeWords for a singleerrorcondition.These changesensure only one errorti_ word wlll be generatedfor a

solid failurecondition.

Instrumentation Delete UHFRF telemetry link. Remove TheCCS is considered operational andand the following equipment: UHFRFAssembly, the backup UHFlink is unnecessary.Communications UMFRF Filter,PCMCoaxialSwitch,CCS

Hybrid Ring, CCS1t4antenna, Add CCSPowerDividerto replacethe HybridRing,

S-IVBand ID PCM signalsrerouted. The S-IVB PCM was removedfromthe CCS

and replacedwith iU PON that had beenroutedto the UHF Transmitter.

Add D68-603 LM RTG Cask DiffuserInletPressure.

DeleteK133-603andK134-603. UHF CoaxlalSwitchmeasurementdeleted.

Structures Add corkInsulatingmaterialto outer IU Withoutcork andwith steelchannels,surfaceand a sheetof vibrationdamping the safety factorat S-[CCECOwas 1,14.

materialin placeof steelchannelsfor The cork and vibration damping compoundthe STI24Mvibrationdamping, increasethis factorto l.SS (l.4Ois

requiredformannedflight).

The DoubleVolumeM/W Accumulator The steel bracketsprovideadequate

mountingbracketswere changedfrom supportfor the increasedloadofaluminumto steel, the newAccumulator. Dynamictestshad

revealedhairlinecracksin the. a

luminumbracke

ts.

Add heavycorematerialin the reglonof Thisgives a highermarginof safetytheWater Accumulatorattachpoints, againstcore crushingunderthe(Location3), attach pads.

Redesignedumbilicalplatedaddedto Internalstiffeningwas addedtoS-IU-505andSubs. increasethestrengthof the plates.

SwingAm testsrevealedexcessive

deflectionof the old platewhen thedisconnect mechanism failed to release

cleanly.

Flight A list of significantlogicchangesadded

Program to the S-IU-503C PrimeLVDCFlightProgramto define the S-IU-SOSFMissionProgram isgivenbelow:

DigitalCoomandSystem(DES)targetandnavigationupdate.

S-IfGuidanceto cutoff.

S-IICECO.

Open loop P/U S-II and S-IVB; differentS-IVBEMR Shift timefor firstand

secondopportunities.PropellantDump andSlingshotmaneuverasseparateTime Base (8) ratherthan,included in TB7.

DCS Command- EnableTB8.

DCSCommand- TD&E Enable.

DCS Command- EnableManeuverA,

GuidanceSwitchover.

Continuousreal-timetelemetry.

B - 4



MPR-SAT-FE-69-7

APPROVAL

S AI'UR N V LA UN CH VE HICLE FL IGHT E VA LUA TIO N R EP OR T

A S -505, A P O LLO I0 M IS S IO N

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 porgrams has been made by the MSFC Security ClassificationOfficer. The highest classification has been determined to be unclassified-

t \ Ef

Stanley L. Fragge

S ecuri ty C las sifica tio n O ffice r

This report has been reviewed and approved for technical accuracy.

George H. McKay, dr. ,jC_irman, SaturnFl-_aluation WorkingGroup/.,,,,, ._-""_"_.7 _ .,_, ,

.,'v , J':"7>'."

Herman K. W eidner

Director, S cience and E ngineering

_ u/Brn_a_gS_ Ma/nger



DISTRIBUTION:

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