· DOCUMENT NO. D5-15795-7 TITLE SATURN V AS-507 LAUNCH VEHICLE OPERATIONAL ABORT AND MALFUNCTIONED FLIGHT ANALYSIS MODEL NO. SATURN V …

D5-15795-7

SATURNV AS-507

LAUNCHVEHICLE

OPERATIONAL

ABORTAND

MALFUNCTIONED

FLIGHTANALYSIS

N73-7169_

Unclas

00/99 16860

JULY30, 1969

i

DOCUMENT NO. D5-15795-7

TITLE SATURN V AS-507 LAUNCH VEHICLE OPERATIONALABORT AND MALFUNCTIONED FLIGHT ANALYSIS

MODEL NO. SATURN V CONTRACT NO. NAS8-5608, SCHEDULE II,

PART IIA, TASK 8.1.4,

AND 8.1.5, DRL 049,

ITEM 169

PREPARED BY :

EMERGENCY DETECTION SYSTEM,

ABORT AND ALTERNATE MISSIONS,

AND

FLIGHT DISPERSIONS AND TRACKING

ORIGINAL RELEASE:

JULY 30, 1969

S. C. KRAUSSE

SECTION MANAGER

FLIGHT SYSTEMS ANALYSIS

ISSUE NO. ISSUED TO

THE _IP_IP_"_'_bI'G COMPANY SPACE DIVISION LAUNCH SYSTEMS BRANCH

REV.SYM

D5-15795-7

REV IS IONS

DESCRIPTION DATE APPROVED

ii

D5-15795-7

ABSTRACT AND LIST OF KEY WORDS

This document contains the abort and malfunctioned flight

analysis conducted for the Saturn V AS-507 launch vehicle.

The effects of various failure modes on S-IC, S-II, and

S-IVB flight are evaluated in terms of abort criteria,mission completion capability, and communications and

tracking surveillance.

Saturn V/Apollo Vehicle

Emergency Detection SystemAbort

Malfunctioned Flight

Flight Dynamics

Control Systems Failures

Propulsion Malfunctions

Accelerometer Failures

Structural Integrity

ControllabilityVehicle Loss

Crew Safety

Acquisition and Loss Data

Vehicle Surveillance

Geometrical Communications Analysis

iii

D5-15795-7

PARAGRAPH

i.i1.21.31.4

2.12.22.32.4

3.13.23.3

4.14.24.34.44.54.64.74.84.94.104.114.124.134.144.154.16

CONTENTS

REVISIONSABSTRACTAND LIST OF KEY WORDSCONTENTSILLUSTRATIONS AND TABLESREFERENCES

SECTION 1 - SUMMARY

MALFUNCTIONDYNAMICSCREWSAFETYPERFORMANCETRACKING AND COMMUNICATIONS

SECTION 2 - EMERGENCYDETECTION SYSTEMDESCRIPTION

EMERGENCYDETECTION SYSTEMDISPLAYSEMERGENCYDETECTION SYSTEMCONTROLSABORT CONTROLSABORTMODES

SECTION 3 - MALFUNCTIONANALYSIS CRITERIA

ABORTCRITERIACREW, VEHICLE, AND MISSION LOSS CRITERIAABORTCUESAND EDS LIMITS

SECTION 4 - MALFUNCTIONEDFLIGHT AND ABORTANALYSIS

SINGLE ENGINE LOSS OF THRUSTDUAL ENGINE LOSS OF THRUSTSINGLE ACTUATORHARDOVERSINGLE ACTUATORINOPERATIVESATURATEDERRORSIGNALSATURATEDRATE SIGNALLOSS OF INERTIAL ATTITUDELOSS OF ATTITUDE COMMANDLOSS OF ATTITUDE ERRORSIGNALLOSS OF ATTITUDE RATE SIGNALACCELEROMETERMALFUNCTIONSPU SYSTEMMALFUNCTIONSLOSS OF ONE APS MODULELOSS OF BOTHAPS MODULESSEQUENCINGAND STAGING MALFUNCTIONSS-II/S-IVB EARLY STAGING

PAGE

iiiiiivvixii

i-i

1-21-21-21-3

2-1

2-12-42-52-5

3-1

3-13-13-4

4-1

4-34-164-334-424-454-514-554-654-674-704-754-904-924-934-944-97

iv

D5-15795-7

CONTENTS (CONTINUED)

PARAGRAPH

SECTION 5 - COMMUNICATIONSANALYSIS FORABORTAND ALTERNATEMISSIONS

5.15.25.3

5.4

5.5

COMMUNICATIONSANALYSIS BACKGROUNDANALYTICAL PROCEDURESAND STUDYLIMITATIONSSURVEILLANCEFOR NONACCELEROMETERMALFUNCTIONSSURVEILLANCEFOR ALTERNATEMISSIONS RESULTINGFROMACCELEROMETERFAILURESS-IVB SECOND-BURNEARLY CUTOFFSURVEILLANCEANALYSIS

PAGE

5-1

5-15-25-4

5-5

5-6

v

D5-15795-7

ILLUSTRATIONS

FIGURE

i-i

1-21-21-31-31-41-41-5

1-5

1-61-61-7

1-7

1-8

1-92-13-14-14-2

4-3

4-44-5

4-6

4-7

4-8

4-9

4-10

4-11

4-12

PAGE

1-4Summary of Malfunctions Requiring Near PadAbortS-IC Malfunction Summary (Sheet 1 of 2) 1-5S-IC Malfunction Summary (Sheet 2 of 2) 1-6S-II Malfunction Summary (Sheet 1 of 2) 1-7S-II Malfunction Summary (Sheet 2 of 2) 1-8S-IVB ist Burn Malfunction Summary (Sheet 1 of 2) 1-9S-IVB ist Burn Malfunction Summary (Sheet 2 of 2) i-i0S-IVB Parking Orbit Coast Malfunction Summary i-ii(Sheet 1 of 2)S-IVB Parking Orbit Coast Malfunction Summary 1-12(Sheet 2 of 2)S-IVB 2nd Burn Malfunction Summary (Sheet 1 of 2) 1-13S-IVB 2nd Burn Malfunction Summary (Sheet 2 of 2) 1-14S-IVB Translunar Coast Malfunction Summary 1-15(Sheet 1 of 2)S-IVB Translunar Coast Malfunction Summary 1-16(Sheet 2 of 2)Summary of Engine-Out and Early Staging 1-17Capability

Summary of Accelerometer Failure Capability 1-18EDS Display and Control Panel 2-2

Abort Lead Time Requirements 3-2

Saturn V Guidance and Control Failure Points 4-2

AS-507 S-IC Engine-Out Guidance x-Freeze 4-5Schedule

Time History of Dynamic Variables for an Early 4-6Engine-Out Malfunction

S-IC Engine-Out Control Capability 4-7

CSM Joint Bending Moment Time H_story Following 4-8Engine Shutdown

Inertial Flight Path Angle Vs Inertial Velocity 4-9Envelope for S-IC Single Engine Failures

Altitude Vs Surface Range for S-IC Single Engine 4-10Failures

Inertial Flight Path Angle Vs Inertial Velocity 4-11Envelope for S-II Single Engine Failures

Altitude Vs Surface Range for S-II Single Engine 4-12Failures

Lead Times for Abort After Early S-IC Engine 4-13Failures

S-IVB First Burn Duration for S-IC Single Lower 4-14

Engine Shutdowns Achieving POI

S-IVB First Burn Duration for S-II Single Lower 4-15

Engine Shutdowns Achieving POI

vi

FIGURE

4-13

4-14

4-15

4-16

4-17

4-18

4-19

4-20

4-21

4-22

4-23

4-24

4-25

4-26

4-27

4-284-29

4-30

4-31

4-32

4-33

4-34

4-35

D5-15795-7

ILLUSTRATIONS (CONTINUED)

S-IC Sequential Engine-Out Capability - EnglnesNo. 2 and No. 3S-IC Sequential Engine-Out Capability - EnglnesNo. 1 and No. 2S-IC Sequential Engine-Out Capability - EnglnesNo. 1 and No. 4S-IC Sequential Engine-Out Capability - EnglnesNo. 3 and No. 4S-IC Sequential Engine-Out Capability - EnglnesNo. 1 and No. 5 or No. 4 and No. 5

S-IC Sequential Engine-Out Capability - EnglnesNo. 1 and No. 3




Mission Completion Capability for S-II No. 1

& 2 Engine Shutdown

Mission Completion Capability for S-II No.

1 & 4 Engine Shutdown







Time History of Dynamic Variables for Engines1 and 4 Out after CECO

Actuator Hardover Control Capability

Time History of Dynamic Variables for Actuator

Hardover at 80 Seconds

Combined Tension Loads Following Actuator

Hardover at 80 Seconds (Sta. 1564)

Time History of Dynamic Variables for Actuator

Hardover Near Gain Change

Combined Tension Loads Following Actuator

Hardover Near Gain Change (Sta. 1760)

S-II Pitch Engine Response to an ActuatorHardover

Lead Times for Abort after an S-IC ActuatorHardover

Divergence Caused by Pitch Actuator Null in

S-IVB Stage

PAGE

4-18

4-19

4-20

4-21

4-22

4-23

4-24

_-25

4-26

4-27

4-28

4-29

4-30

4-31

4-32

4-35

4-36

4-37

4-38

4-39

4-40

4-41

4-43

vii

FIGURE

4-36

4-37

4-38

4-39

4-40

4-41

4-42

4-43

4-444-45

4-46

4-47

4-48

4-49

4-50

4-514-52

4-53

4-54

4-554-56

4-57

4-58

4-59

D5-15795-7


PAGE

Abort Lead Time for Yaw Actuator Null in S-IVB 4-44FlightTime History of Dynamic Variables for Saturated 4-47Error Signal (S-IC)Tension Load at Station 2832 Aft (Saturated 4-48Error Signal)Vehicle Pitch Rate History Following Saturated 4-49Error Signal in S-II and S-IVBAbort Lead Times for Aborts Following a Saturated 4-50Attitude Error Malfunction in S-ICTime History of Dynamic Variables for Saturated 4-52Rate Signal (S-ICTension Load at Station 1760 (Saturated Rate 4-53Signal)Vehicle Pitch Rate History Following Saturated 4-54Rate Signal in S-II and S-IVBGuidance Switchover Sequence 4-57Positive Pitch Platform Failure Capability With 4-58Spacecraft BackupNegative Pitch Platform Failure Capability With 4-59

Spacecraft Backup

Yaw Platform Failure Capability With Spacecraft 4-60Backup

Minimum Pitch Attitude Errors At Switchover 4-61

for Platform Failures that Cause Loss of Control

Minimum Yaw Attitude Errors At Switchover for 4-62

Platform Failures that Cause Loss of Control

Maximum Tolerable Drift Rate with 75 NMI Orbit 4-63

Capability

Lead Times for Abort After S-IC Loss of Platform 4-64

Time History of Dynamic Variables for Loss of 4-68

Attitude Error (S-IC)

Tension Loads at Sta. 1564 (Loss of Attitude 4-69

Error)


Attitude Rate (S-IC)

Tension Load at Sta. 1564 (Loss of Attitude Rate) 4-72


Attitude Rate (S-II)

S-IVB Roll Dynamics Following Loss of Roll Rate 4-74

Signal During Second Burn

Apogee/Perigee Altitude for X-Axis Accelerometer 4-79Failures

Apogee/Perigee Altitude for Z-Axis Accelerometer 4-80Failures

viii

FIGURE

4-60

4-61

4-62

4-63

4-64

4-65

4-66

4-67

4-68

4-69

5-1

5-4

5-5

5-6

5-7

5-8

5-9

5-10

D5-15795-7

ILLUSTRATIONS (Continued)

Yaw Attitude History for Y-Axis AccelerometerFailure at 2 Seconds (Boost to Insertion)Yaw Attitude History for Y-Axis AccelerometerFailure at 2 Seconds from First Motion(Boost to TLI)

Velocity Error at TLI in Failed Axis forX-Accelerometer Failures in S-IVB Second BurnVelocity Error at TLI in Failed Axis forY-Accelerometer Failures in S-IVB Second BurnVelocity Error at TLI in Failed Axis forZ-Accelerometer Failures in S-IVB Second BurnCommanded Pitch Attitude Envelope for Accelero-meter Failure During First Opportunity BurnCommanded Yaw Attitude Envelope for Acceler-ometer Failures During First Opportunity BurnVelocity Error in Failed Axis for X-Acceler-ometer Failure for First OpportunityVelocity Error in Failed Axis for Z-Acceler-ometer Failure for First OpportunityS-IVB Burn Duration Vs Time of S-II/S-IVBEarly StagingEastern Test Range Minimum Altitude/SurfaceRange 0-Degree Elevation Angle Profile forContinuous Tracking Coverage Based on Mila,

Grand Turk, Grand Bahama, Antigua, and BermudaGrand Stations

Nominal First Parking Orbit Groundtrack with

0-Degree Elevation Angle Visibility Envelopes

Nominal Second Parking Orbit Groundtrack

with 0-Degree Elevation Angle Visibility

Envelopes

Nominal Third Parking Orbit Groundtrack with

0-Degree Elevation Angle Visibility EnvelopesTracking Delta Times for NonaccelerometerMalfunctions

Nominal Surveillance Above 0 Degree for FirstOrbit

Nominal Surveillance Above 0 Degree for SecondOrbit

Nominal Surveillance Above 0 Degree for ThirdOrbit

0-Degree Acquisition and Loss History forFirst Orbit for X-Accelerometer Malfunctions

for 78.051 ° Launch Azimuth

0-Degree Acquisition and Loss History for SecondOrbit for X-Accelerometer Malfunctions for

78.051 ° Launch Azimuth

PAGE

4-81

4-82

4-83

4-84

4-85

4-86

4-87

4-88

4-89

4-98

5-10

5-12

5-13

5-14

5-15

5-16

5-17

5-18

5-19

5-20

ix

FIGURE

5-11

5-12

5-13

5-14

5-15

5-16

5-17

5-18

5-19

5-20

5-21

5-22

5-23

D5-15795-7


0-Degree Acquisition and Loss History for ThirdOrbit for X-Accelerometer Malfunctions for78.051 ° Launch Azimuth0-Degree Acquisition and Loss History for FourthOrbit for X-Accelerometer Malfunctions for78.051 ° Launch AzimuthAcquisition Azimuth for X-Accelerometer Mal-functions Above 0-Degree Elevation for 78.051 °Launch Azimuth

0-Degree Acquisition and Loss History for FirstOrbit for Z-Accelerometer Malfunctions for78.051 ° Launch Azimuth

0-Degree Acquisition and Loss History for Second

Orbit for Z-Accelerometer Malfunctions for78.051 ° Launch Azimuth

0-Degree Acquisition and Loss History for Third

Orbit for Z-Accelerometer Malfunctions for


0-Degree Acquisition and Loss History for Fourth

Orbit for Z-Accelerometer Malfunctions for


Acquisition Azimuth for Z-Accelerometer Mal-

functions Above 0-Degree Elevation for 78.051 °Launch Azimuth

Nominal Surveillance Above a 0-Degree Elevation

Angle for Fourth Parking Orbit (No S-IVB SecondBurn)

Nominal Surveillance Above a 0-Degree Elevation

Angle for Fifth Parking Orbit (No S-IVB SecondBurn)

Nominal Surveillance Above a 0-Degree ElevationAngle for Sixth Parking Orbit (No S-IVB Second

Burn)

S-IVB Stage Early Second Cutoff Surveillance

Above 0-Degree Elevation for 78.051 ° LaunchAzimuth

S-IVB Stage Early Second Cutoff Acquisition

Azimuth Above 0-Degree Elevation for 78.051 °Launch Azimuth

PAGE

5-21

5-22

5-23

5-24

5-25

5-26

5-27

5-28

5-29

5-30

5-31

5-32

5-33

x

D5-15795-7

TABLE

3-I

4-I

4-II

4-III

5-I

5-II

5-III

5-IV

TABLES

Manual Abort Cues and EDS Limits

TLI Capability for PU Malfunctions

Abort Logic and Mission Capability forSequencing and Staging Malfunctions

S-II/S-IVB Early Staging Conditions for

African Continent ImpactsSurveillance Network Used for AS-507 G

Mission Communications Analysis forAbort and Alternate Missions

Summary of the 0-Degree Surveillance for

the Nominal Vehicle During EPO

Alternate Missions Achieving a NominalParking Orbit

Summary of the 0-Degree Surveillance for

a Y-Accelerometer Malfunction at 2 Seconds

PAGE

3-5

4-91

4-96

4-97

5-8

5-9

5-11

5-34

xi

D5-15795-7

IQ

o

o

o

1

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o

•

REFERENCES

Boeing Document D5-15795-6, "Saturn V AS-506 Launch

Vehicle Operational Abort and Malfunctioned Flight

" dated June 3, 1969Analysis,

Boeing Coordination Sheet EDS-H-256, "EDS Criticality

Determination and Bar Chart Summaries," dated

April 29, 1969.

MSFC Memorandum S&E-ASTN-DIR-69-150, "Saturn V

Overall Failure Mode Criticality and Probabilityof Occurrence," dated March 27, 1969.

MSFC Memorandum S&E-ASTN-STA-69-2, "Saturn V Overall

Failure Mode Criticality and Probability of Occurrence -

" dated June 25, 1969AS-506,

Boeing Document D5-15795-5A, "Saturn V AS-505 Launch

Vehicle Operational Abort and Malfunctioned Flight

" dated April 25, 1969Analysis,

MSFC Document I-MO-3-69, "Final Flight Mission Rule

Input Document," dated May 16, 1966. (AS-506)

Boeing Memorandum 5-9600-H-280, "SSR-249, Spacecraft

Backup for Saturn V Inertial Platform, AS-505,"

dated May 14, 1969.

MSC Memorandum ET 253/6810-170-B, "Required Lead

Times, Saturn V," dated October 3, 1968.

k

xii

D5-15795-7

SECTION 1

SUMMARY

This document contains theresults of malfunctioned flightand abort analysis for the AS-507 G mission. The G mission

is a lunar landing mission with launch scheduled for the

September-October launch window.

Analysis results reported in this document include the

following:

a.

Do

Co

Evaluation of the operational effectiveness of the

Emergency Detection System (EDS) automatic and manual

abort limits to provide crew safety and minimize falseaborts

Vehicle dynamic and structural load responses, regions

of controllability, and trajectory envelopes resultingfrom vehicle malfunction

Vehicle capability to achieve the performance plateaus

defined by launch vehicle mission requirements - earth

parking orbit, translunar injection, and a 75 nautical

mile contingency orbit

d. Verification of LVDC flight program presettings for

malfunctioned flight

eo

An evaluation of the surveillance network's capability

to provide adequate tracking and communications duringmalfunctioned flight.

The AS-507 G mission analysis is similar to that for the

AS-506 G mission, reported in Reference i, with thefollowing major differences:

a.

b.

Co

do

The EDS rate limit between 120 seconds and S-IC OBECO

is changed from 4 degrees per second to 10 degreesper second.

The setting for the overrate light between 120 seconds

and 134 seconds is changed from 4 degrees per second toi0 degrees per second.

The effects of the S-IC Automatic Abort Cutoff (AACO)

logic following dual adjacent engines out after CECOare evaluated in detail.

The analysis is expanded to evaluate the effects of

single-axis platform failures for various failure rates

and accelerations.

i-i

D5-15795-7

SUMMARY (Continued)

e • The effects of accelerometer failures coupled with thrust

misalignments and failures across various launch windows

are evaluated.

Significant results derived from this analysis are summarized

in the following paragraphs.

i.i MALFUNCTION DYNAMICS

Malfunction dynamics for the AS-507 are essentially the same

as for AS-506, in spite of increased winds. Times during

which loss of control occurs are approximately the same for

AS-507 as for AS-506. Control is maintained for adjacent

simultaneous S-IC engine failures occuring after 135 seconds.

This is due to all engines being shut down when any two adja-

cent control engines are out after 135 seconds• Loss of con-

trol in S-II, however, can result from sequential S-IC

adjacent engine failures where the second engine fails before

approximately 150 seconds.

Spacecraft loads after engine-out are slightly increased for

AS-507 because of higher winds but remain below structural

limits for all flight times.

A summary of vehicle controllability for all types of mal-

functions in all stages is shown in Figures i-i through 1-8.

1.2 CREW SAFETY

Periods of possible crew loss for AS-507 are essentially the

same as for AS-506.

These periods occur only for short time intervals during S-IC

flight and are the result of a single engine out, an actuator

hardover, or a saturated error signal. No other malfunctions

considered cause crew loss.

The crew loss criticalities for these three malfunctions,

as well as the flight times during which crew loss can occur,

are shown in Figure 1-2. Probability of crew loss due to

catastrophic S-IC engine failure remains unchanged at61 X 10 -°.

1.3 PERFORMANCE

Predicted malfunctioned flight performance summarized in

Figure 1-8 for AS-507 indicates generally improved POI/TLI

capability over AS-506 for propulsion failures. TLI may

1-2

D5-15795-7

SUMMARY (Continued)

be achieved with an S-IC single engine out approximately 15seconds prior to that predicted for AS-506. TLI may be achievedwith an S-II single engine out 30 seconds prior to that pre-dicted for AS-506. Improvement for other propulsion failuresranges from 5 to 15 seconds. Parking orbit capability existsfor any S-IC adjacent control engine failure occuring afterapproximately 150 seconds. Capability to achieve an orbitgreater than 75 NMI perigee for accelerometer failures is un-changed from that predicted for AS-506, except for X-axisaccelerometer failures with 3-sigma high and low performingvehicles. The earliest 3-sigma high failure time to 75 NMIperigee orbit has been extended from ii0 seconds from firstmotion for AS-506 to 135 seconds for AS-507. The 3-sigma lowtime has been extended from 0.0 second to 30 seconds.

A summary of AS-507 POI/TLI capability for engine out and

early staging is given in Figure 1-8 and for accelerometer

failures in Figure 1-9.

Improvement in propulsion malfunction performance is due to

the steeper trajectory profile for AS-507 during S-IC and the

first portion of S-II burn; this results in a higher altitude

for the same failure time. The AS-507 into-orbit pitch pro-

file differs from the AS-506 pitch profile and results in

lower perigees for X-accelerometer failures with the presentpresettings.

1.4 TRACKING AND COMMUNICATIONS

The communications and tracking analysis indicates that the

geometrical coverage for malfunctioned flight is comparable

to that for nominal flight.

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SECTION 2

EMERGENCY DETECTION SYSTEM DESCRIPTION

The Saturn V Emergency Detection System (EDS) is designed to

provide crew safety in the event of malfunctioned flight. The

system monitors critical flight parameters and furnishes advance

warning of impending emergency conditions. If abort is required,

it is initiated either automatically or manually depending onthe nature and time of the malfunction.

In the following paragraphs, the EDS system is described in

terms of EDS displays, EDS controls, abort controls, and abortmodes.

2.1 EMERGENCY DETECTION SYSTEM DISPLAYS

The EDS displays are selected to present parameters which

indicate failures leading to vehicle abort. Automatic abort

parameters are implemented triple redundant, voted two-out-of-

three, to preclude single point hardware or sensing failures

causing an inadvertant abort. Manual abort parameters are

implemented with redundant sensing and displays to providehighly reliable indications to the crew.

Displays are designed to provide onboard detection capabilityfor rapid rate malfunctions which may require abort. Pilot

abort action must be based on two separate but related abort

cues. These cues may be derived from the EDS displays, groundinformation, physiological cues, or any combination of two

valid cues. The EDS displays and controls referred to through-

out this document are shown in Figure 2-1. As each is discussed,it is identified by use of grid designators listed on the borderof the figure.

2.1.1 Flight Director Attitude Indicator (FDAI)

There are two Flight Director Attitude Indicators, each of

which provides a display of Euler attitude, attitude errors

and angular rates. These displays are active at liftoff and

remain active throughout the mission, except that attitude

errors are not displayed during S-II and S-IVB flight. The

FDAI's are used to monitor normal launch vehicle guidance and

control events. The pilot's FDAI is shown in Figure 2-1, I-9.

The FDAI ball displays the Euler attitudes, the needle type

pointers across the face of the ball indicate attitude errors,

and the triangular pointers around the periphery of the balldisplay angular rates.

2-1

D5-15795-7

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1'0 1"I 12 1'3

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EDSDISPLAYArJDCONTROLPAIIEL

2-2

D5-15795-7

2.1.1 (Continued)

Signal inputs to the FDAI's are switch selectable and can come

from a number of different sources in the spacecraft. This

flexibility and redundancy provide the required attitude and

error backup display capability.

2.1.2 LV ENGINES Lights

Each of the five LV ENGINES lights shown in Figure 2-1, J-12,

represents the respective numbered engine on the operating

stage. A light ON indicates its corresponding engine is

operating below a nominal thrust level (90 percent on F-I

engines and 65 percent of J-2 engines). During staging all

lights are turned OFF momentarily to indicate physical

separation has occurred.

2.1.3 LV RATE Light

The LV RATE light (Figure 2-1, 1-12) when ON, is the primary

cue from the launch vehicle that the following preset over-

rate settings have been exceeded:

Pitch/Yaw

Roll

4.0 (±0.5) deg/sec Liftoff to automatic

abort deactivation

(120 seconds)

9.2 (±0.8) deg/sec Automatic abort

deactivation to

S-IVB cutoff

20.0 (±0.5) deg/sec Liftoff to S-IVB

cutoff.

In the event the LV GUID light is illuminated during the auto-

matic abort phase, the LV RATE light will be illuminated as a

redundant indicator.

2.1.4 LV GUID Light

Launch vehicle attitudes are measured and provided to the

launch vehicle digital computer every 40 milliseconds. The

computer checks the attitude for reasonableness. If the

reasonableness tests fail, the attitude error signals to the

flight control computer are frozen and the LV GUID light,

shown in Figure 2-1, 1-12, is illuminated.

2-3

D5-15795-7

2. i. 5 ABORT Light

The ABORT light, shown in Figure 2-1, G-12, may be illuminatedby ground command from the Flight Director, the Mission Control

Center (MCC) Booster Systems Engineer, the Flight Dynamics

Officer (FDO), the Complex 39 Launch Operations Manager (until

tower clearance +i0 seconds), or in conjunction with range

safety booster engine cutoff.

2.1.6 Angle-of-Attack Meter

The angle-of-attack (qu) meter, shown in Figure 2-1, L-8, is

time shared with the Service Propulsion System (SPS) chamber

pressure. The q_ display is a pitch and yaw vector summed

angle-of-attack/dynamic pressure product. It is expressed in

percentage of total pressure for predicted launch vehicle

breakup (abort limit equals i00 percent). It is effective as

an abort parameter only during the high q flight region. During

other portions of boost through the atmosphere, the qe meter

provides trend information on launch vehicle flight performance

and provides a secondary cue for slow-rate guidance and controlmalfunctions.

2.1.7 Accelerometer

The accelerometer, shown in Figure 2-1, H-6, indicates longi-

tudinal acceleration/deceleration in G's. It provides a

secondary cue for certain engine failures and is a gross indi-

cator of launch vehicle performance.

2.1.8 Event Timer

The event timer, shown in Figure 2-1, H-12, is a digital clock

which displays time from liftoff. It is a critical display

because it is the primary cue for the transition of abort modes,

manual sequenced events, monitoring roll and pitch program,

staging, and S-IVB insertion cutoff.

The event timer is reset to zero automatically with abort initia-

tion.

2.2 EMERGENCY DETECTION SYSTEM CONTROLS

The main EDS control switches are the EDS, the 2 ENG OUT, and

the LV RATES. They are two-position (AUTO and OFF) toggle

switches which are placed in the AUTO position prior to

liftoff. When all three switches are in the AUTO position,

automatic abort is initiated if:

a. A LV structural failure occurs between the Instrument Unit

and Command Service Module.

b. Two or more S-IC engines drop below 90 percent thrust.

2-4

D5-15795-7

k

2.2 (Continued)

c. LV rates exceed the EDS limits.

While these switches can be manually disabled at any time by

placing them in the OFF position, the normal procedure requires

disabling at 120 seconds. They are automatically disabled by

the LV sequencer just prior to center engine cutoff.

2.3 ABORT CONTROLS

The translational controller (T-Handle) mounted on the left

arm of the commander's couch is used to initiate abort. A

manual Launch Escape System (LES) abort sequence is initiated

by rotating the T-Handle fully counterclockwise. This sends

redundant engine cutoff commands to the LV, initiates Command

Module/Service Module separation, fires the LES motors, resets

the spacecraft sequencer and initiates the post abort sequence.

(Engine cutoff from the spacecraft is inhibited during thefirst 30 seconds of flight).

For a manually initiated SPS abort after the Launch EscapeTower has been jettisoned, counterclockwise rotation of the

T-Handle commands LV cutoff, resets the spacecraft sequencerand initiates the Command Service Module/Launch Vehicle

separation sequence. However, returning the T-Handle to

neutral before 3 seconds expires results in only a LV cutoff

rather than the full abort sequence.

2.4 ABORT MODES

Aborts during boost are performed using either the LES or theSPS.

2.4.1 LES Aborts (Modes IA, IB, and iC)

The LES consists of a solid propellant launch escape motor

used to propel the CM a safe distance from the launch vehicle,

a tower jettison motor, and a canard subsystem. LES abortmodes are as follows:

a. Mode IA: Low Altitude Mode (Pad to 42 Seconds)

In Mode IA a pitch control motor, mounted normal to

the launch escape motor, propels the vehicle downrange

to ensure water landing and escape from the "fireball."

The automatic sequence of major events following abortinitiation is as follows:

2-5

D5-15795-7

2.4.1 (Continued)

Time Event

00:00

00:05

00:Ii

00:14.4

00:16

00:28

Abort, SM Reaction Control System

(RCS) Oxidizer Rapid Dump, Launch

Escape & Pitch Control Motors Fire

RCS Fuel Rapid Dump

Canards Deploy

Apex Cover Jettison

Drogue Deploy

Main Chute Deploy

The automatic sequence can be prevented, interrupted, or

replaced by crew action.

b. Mode IB: Medium Altitude Mode (42 Seconds to i00,000 Feet)

Mode IB is essentially the same as Mode IA with the ex-

ception of deleting rapid RCS propellant dump and PC motor

fire. The canard subsystem is designed specifically for

this altitude region to initiate a tumble in the pitch

plane. Upon closure of barometric switches at 24,000

feet, the tower is jettisoned. The main parachutes are

automatically deployed at 10,000 feet.

C. Mode IC: High Altitude Mode (100,000 Feet to Tower

Jettison)

During Mode iC the launch vehicle is above the atmosphere

and therefore the canard subsystem cannot be used to

induce a pitch rate to the escape vehicle. If the launch

vehicle is stable at abort, the LET is manually jettisoned

and the CM oriented to the reentry attitude. This method

requires a functioning attitude reference system.

With a failed attitude reference system the alternate

method is to introduce a 5 degree per second pitch rate

using the attitude control thrusters. The CM/tower com-

bination will then stabilize blunt end forward as in

Mode lB. The LES then deploys the parachutes at the properaltitude.

2.4.2 SPS Aborts (Modes II, III, and IV)

The SPS aborts utilize the Service Module SPS engine to propel

the CSM combination away from the launch vehicle, maneuver for

reentry to a planned landing area, or boost into a contingencyorbit. The SPS abort modes are as follows:

2-6

D5-15795-7

2.4.2 (Continued)

a. Mode II

The SM Reaction Control System engines are used to propelthe CSM away from the launch vehicle unless the vehicle isin danger of exploding or excessive tumble rates are pre-sent at LV/CSM separation. In these two cases the SPSengine would be used due to greater AV and attitude con-trol capability. When at a safe distance, the CM isseparated from the SM and maneuvered to a reentry attitude.

b. Mode III

The SPS engine is used to slow £he CSM combination so asto land at a predetermined point in the Atlantic Ocean.CM/SM separation then occurs and normal reentry proceduresfollow.

c. Mode IV

The SPS engine can be used to make up for a deficiency in

insertion velocity up to approximately 3000 feet per

second. This is accomplished by holding the CSM in an

inertial attitude and applying the needed AV with the

SPS to acquire the acceptable orbital velocity.

2-7

D5-15795-7

THIS PAGE INTENTIONALLY LEFT BLANK.

2-8

D5-15795-7

SECTION 3

MALFUNCTION ANALYSIS CRITERIA

Malfunction analysis criteria consists of:

at

b.

c.

Abort criteria

Crew, vehicle, and mission loss criteria

Abort cues and EDS limits

3.1 ABORT CRITERIA

Abort criteria are those criteria which are used to establish

the need for abort in the event of a vehicle malfunction. Thecriteria are:

a.

b,

C.

d.

e.

f.

g.h.

Launch platform interference, pad fallback, and towercollision.

Controllability

Structural capability

Range safety limits

Staging limits

Spacecraft heating limit

Platform yaw gimbal limit

Safe abort limits

Conditions for a safe abort must satisfy the following re-

quirements and limits:

a.

b.

c.

d.

e°

f.

3.2

Water impact

Abort lead time requirements (Figure 3-1)LEV u-limit

16g reentry limiti00 second free fall limit

Spacecraft platform tumble limit (90 degrees yaw)

CREW, VEHICLE, AND MISSION LOSS CRITERIA

3.2.1 Crew Loss

Crew loss is assumed to occur if the abort lead time is less than

the required lead time, or if 90 degrees yaw attitude occurs dur-

ing abort. Abort lead time requirements during S-IC flight are

taken from Reference 8 and shown in Figure 3-1.

Crew loss factors (8's) are calculated as follows:

(_TLoss) (PLoss)

STAGE FLIGHT TIME

3-1

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3-2

D5-15795-7

3.2.1 (Continued)

where: _TLoss

PLOSS

= time interval in which crew loss occurs

due to a particular malfunction

= probability that the particular malfunction

will cause crew loss; i.e., if, considering

single engine failures in S-IC, any control

engine failure causes crew loss, then

P S = 4/5; or if only a particular engine

c_es loss of control, then PLOSS = 1/5

In case of crew loss 8's, it is assumed that the probability

of an explosion within one second following breakup _s 0.2.

Criticality numbers are calculated as follows:

CN = B • UN

where UN (unreliability) is defined as the probable number of

occurrences of a failure mode per one million flights.

A detailed explanation of loss factor determination is given

in Reference 2. Unreliability values used in this analysisare given in References 3 and 4.

3.2.2 Vehicle Loss

Vehicle loss is assumed to occur if the vehicle or spacecraft

structural capability is exceeded before the vehicle or payload

achieves at least a 75 NMI perigee orbit. Situations causingstructural failure are:

a.

Do

Vehicle collision with a solid object, e.g., liftoff

interference with launch platform and holddown posts,

tower collision, pad area fallback, and collision be-

tween separating stages following staging.

Excessive aerodynamic forces due to loss of control or

trajectory deviations which lead to atmospheric re-entry.

C,

Structural dynamic response following a malfunction andabort cutoff.

Vehicle loss factors are calculated in the same manner ascrew loss factors.

3-3

D5-15795-7

3.2.3 Mission Loss

Launch vehicle primary mission loss is assumed if, at TLIplus 3 hours, a AV correction greater than 800 feet per secondis required.

Mission loss factors are calculated in the same manner ascrew loss factors.

3.3 ABORT CUESAND EDS LIMITS

Automatic abort occurs during the first 120 seconds of flightwhen:

a. A vehicle overrate is measured by two of the three EDS

rate gyros in each axis. An overrate is 4 degrees per

second in pitch or yaw, or 20 degrees per second in roll.

b. Two engines are below 90 percent thrust as indicated by

two of the three "Thrust-OK" pressure switches on each

engine.

C. A structural failure occurs between the Instrument Unit

and the Command Module as indicated by two out of threestructural wires.

Manual abort will be initiated by the crew using the abort

cues and EDS limits shown in Table 3-I. Two independent

cues measured and indicated by separate systems are necessaryto prevent false abort.

3-4

TABLE 3-I

DS-I 5795-7

MANUAL ABORT CUES AND EDS LIMITS

STAGE

S-IC

S-II&

S-IVB

FLIGHTTIME(SEC)

O<t<50

50<t<120

120<t<OBECO

AllTimes

ABORT CUE

Engine Status LightsVoice RequestAbort Request Light

Engine Status LightsAttitude Error

Q-Ball Ap

Engine Status LightL/V Rate LightS/C Rate Indicator:

RollPitch or Yaw

Attitude Error

S/C Rate Indicator:RollPitch or Yaw

FDO DisplayAbort Request LightL/V Rate LightEngine Status LightsAttitude DeviationYaw Attitude

Voice Request

EDS LIMIT

-+5 deg3.2 PSlD (100%)

-+20 deg/sec

-+I0 deg/sec

±5 deg*

±20 deg/sec

±lO deg/secLimit Exceeded

-+20 deg

±45 deg

*Recommended EDS limit for dual engine out malfunction

occurring after CECO is -+20 degrees.

3-5

D5-15795-7

THIS PAGE INTENTIONALLY LEFT BLANK.

3-6

D5-15795-7

SECTION 4

MALFUNCTIONED FLIGHT AND ABORT ANALYSIS

The objectives of this analysis are to:

a. Determine that the Emergency Detection System protects

the crew following vehicle malfunctions

Do Determine the mission completion capability following

vehicle malfunctions.

The scope of the analysis is limited to the evaluation of

malfunctioned flight for a G mission with a September 13, 1969,

78.051 degree launch azimuth, and first opportunity translunar

injection.

The points of failure for the various malfunctions considered

in this analysis are shown in Figure 4-1.

Except for the effects of accelerometer malfunctions, all

effects of malfunctions are evaluated for a reference vehicle.

Accelerometer malfunction effects are evaluated for ±3-sigmavehicles.

The AS-507 S-IC pitch polynomial is biased for the average

50 percentile September/October/November winds. For malfunctions

where winds have a significant effect, the vehicle is flown

in design winds with magnitudes based on the maximum 95

percentile September/October wind. The gust is phased with

the malfunction to establish a worst case. For malfunctions

where winds do not have a significant effect, the average 50

percentile September/October wind from 258 degrees is used.

4-1

D5-I 5795-7

[ INERTIAL SUBSYSTEM'-]

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I ATTITUDERATE

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C

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---ITHRUST VECTOR Q

INGLE AND DUAL ENGINE LOSS OFHRUST

F _ m mGUIDANCE & NAVIGATION SUBSYSTEMCONNAND ANGLES X

II _1 CO"PARE_ TO_

""-[ AND COMPUTE *ILYDC/LVDA

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,qI'---ATTITUDEERROR

_1 COMBINE & FILTER INPUTS

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BRc' BPc, & BYc

TVC FLUID 1POWER SUBSYSTEM I

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1' I

IENGINE IACTUATORS I

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SINGLE ACTUATOR HARDOVER

SINGLE ACTUATOR TO NULL

C)_T,.,RA',_OCO,_TRO,_,ONALC._'-O_O'A'T*T'.'O__'_RO,_S,O,',_,

C) LOSS OF INERTIAL ATTITUDE i (_ 'ENGINE I

I SUgSYSTEM I)LOSS OF ATTITUDE RATE SIGNAL i tE,GINEs

)LOSS OF ATTITUDE COMMAND SIGNAL

FIGURE 4-i SATURrJV GUIDANCEAND CONTROL FAILURE POINTS

4-2

D5-15795-7

4.1 SINGLE ENGINE LOSS OF THRUST

4.1.1 Malfunction Description

Single engine loss of thrust malfunctions are categorized asfollows:

a. Thrust loss following a shutdown command initiated by

the thrust OK pressure switches (TOPS). A TOPS shutdown

occurs when some malfunction causes propellant inlet

pressure, and ultimately thrust, to drop below 90 percentof rated thrust.

hl Thrust loss resulting from an inadvertant shutdown

command initiated by an electrical malfunction.

C. Sudden thrust loss resulting from a catastrophic failure

such as engine explosion.

Failure of a J-2 engine to start during S-II and S-IVB flight

is also considered in this analysis.

Thrust decay characteristics resulting from TOPS dropout are

shown in Reference 5 . TOPS dropout is inhibited until

TBI + 14 seconds, after which TOPS dropout occurs automatically

when the thrust decreases below 90 percent of rated thrust.

The S-IC engine-out guidance x-freeze schedule used in this

study is shown in Figure 4-2.

4.1.2 Malfunction Dynamics

Any engine out prior to 0.2 second results in pad fallback.

Any engine out between 0.2 and 0.9 second results in the

vehicle colliding with the holddown posts. Tower collision

occurs for a Number 1 or 2 engine out prior to 5.7 seconds.

Number 4 engine out before 3.5 seconds results in loss of

control in the late high-q region due to vehicle aerodynamic

instability and eventual loss of control authority. The

max-q region occurs between ii0 and 130 seconds for these

very early failures. Figure 4-3 shows a time history of

dynamics for a typical early single engine failure.

Figure 4-4 shows that for single engine-out malfunctions in

the high-q region, from approximately 60 to 95 seconds, con-

trollability exists for all 95 percentile September/October

winds. Staging criteria are met for all engine_ou_ cases thatmaintain control to cutoff.

4-3

D5-15795-7

4. i. 2 (Continued)

Structural capability is not exceeded at the Command Service

Module joint following TOPS shutdown. Figure 4-5 shows a

typical bending moment history following TOPS shutdown.

Figure 4-6 shows inertial flight path angle versus inertial

velocity envelope for S-IC single engine failures. In

deriving envelopes, engine failures are simulated with a

3-sigma performance dispersion for the failed stage. The

i00 second free-flight-to-reentry limit is violated for

some S-IC engine shutdown times. However, in most cases

this violation occurs either with an altitude greater than

nominal, or prior to launch escape tower jettison. Figure 4-7

shows the altitude versus surface range envelope for S-ICsingle engine failures.

S-II single engine-out failures do not require abort. S-II

envelopes of flight path angle versus inertial velocity, and

altitude versus surface range for this malfunction are shown

in Figures 4-8 and 4-9 , respectively.

S-IVB engine failure requires staging to the Command ServiceModule for abort.

4.1.3 EDS Effectiveness and Crew Safety

Since there are no effective abort cues for early failures

resulting in pad fallback, holddown post collision or tower

collision, it is assumed that these failures result increw loss.

The l0 degree per second manual cue provides positive abort

lead times for late high-q aborts resulting from early

failures. These abort lead times are shown in Figure 4-10.

False automatic aborts on the 4 degree per second rate limit

can occur for engines 2 or 3 out in the 75 to 80 second

region.

4.1.4 Mission Completion Capability

Figure 1-8 summarizes orbital capability for S-IC and S-IIsingle engine-out failures. S-IVB first burn duration as a

function of S-IC and S-II malfunction time is shown in Figures

4-11 and 4-12, respectively. For S-IVB failure, Command

Service Module parking orbit insertion is possible provided

the failure occurs after 605 seconds flight time, and assuming

Service Propulsion System _V capability to be 500 meters persecond.

4-4

D5-15795-7

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4.2 DUAL ENGINE LOSS OF THRUST


Loss of thrust engine failures are described in Section 4.1.

When loss of thrust on two adjacent control engines is detected

during S-IC flight, after Automatic Abort Cutoff (AACO) is

enabled, the remaining engines are automatically shut down.


Ability to maintain control after failure of two engines in

S-IC or S-II flight depends on which engines have failed and

the times at which they fail. Figures 4-13 through 4-26 show

the failure times which result in loss of control for each

pair of engines. Figure 1-8 summarizes these times for simul-

taneous engine failures.

The regions shown in Figures 4-13 through 4-16 for adjacent

control engines include the effects of early S-IC shutdown.Significant events during the time from 134.8 seconds to S-IIignition are:

EVENT FLIGHT TIME

(SECONDS)

AACO Enable 134.8

CECO & Timebase 2 Set 135.0

Timebase 3 Enable 152.0

OECO & Timebase 3 Set (Nominal) 159.9

S-II at 90% Thrust (Nominal) 164.3

Note that timebase 3 cannot be set until it is enabled 17.0

seconds after CECO. If shutdown occurs at 134.8, there are

21.6 seconds of coast between S-IC shutdown and the time S-II

reaches 90% of full thrust. This extended coast period,

coupled with a significant attitude rate, causes loss of con-

trol. Figure 4-27 shows typical dynamics for two adjacentcontrol engines out after CECO.

No loss of control occurs for a single S-IC engine out

followed by a single S-II engine out.


During S-IC flight automatic abort is initiated if the thrust

OK pressure switches on two engines drop out. This dual

engine out automatic abort may be inhibited manually by the

crew, but is inhibited automatically just prior to CECO.

4-16

D5-15795-7

L 4.2.3 (Continued)

The recommended time to inhibit automatic abort is 120 seconds

in nominal flight because failure of the CSM joint can cause

crew loss for simultaneous failures prior to this time.

After 120 seconds, manual abort cues are:

a,

b.

C.

Engine-out lights

Ten degrees/second overrate light and spacecraft rateindicator

Abort request light

An additional abort cue of +20 degrees attitude error is

recommended for two control-engines out after CECO. With this

cue, no crew loss occurs for this malfunction.


Mission completion capability for two engines out is shown in

Figure 1-8 and in Figures 4-13 through 4-27.

For adjacent S-IC control engines out after AACO enable, the

vehicle fails to reach a 75 nautical mile orbit if shutdown

occurs prior to 149 seconds. Loss of thrust acceleration

causes failure of the accelerometer reasonableness test.

The boost navigator uses backup F/M tables to calculate velo-

city resulting in a position error. Three failures of the

accelerometer reasonableness test cause the orbit to have

a perigee less than 75 nautical miles. Therefore, if the

second engine fails prior to 149 seconds, the orbit is unsafe.

A possible method of eliminating the long coast is to begin

timebase 3 at shutdown. This eliminates most loss of control

cases and permits a safe orbit for all cases that do notlose control.

4-17

D5-15795-7

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TIME HISTORY OF DYNAMICVARIABLES FOR

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4-32

D5-15795-7

4.3 ACTUATOR HARDOVER


Actuator hardover is any failure that causes an engine actuatorto either extend or retract to its limit.


Analysis shows that during S-IC flight, an actuator fully

extended (engine deflected inboard) gives worst case results

due to the S-IC outboard engine cant of two degrees.

An actuator hardover near the pad does not cause pad fallback,but any actuator hardover before 0.7 seconds results in an

engine bell colliding with the holddown posts. An actuator

hardover in the positive yaw direction before 3.2 secondsresults in tower collision.

Figure 4-28 shows control capability for an actuator hardover

in the high-q region. An actuator hardover causes loss of

control in the 95 percentile wind rose between 68 and 88

seconds (high-q region). Figure 4-29 shows typical dynamicsfor an actuator hardover at 80 seconds. Loss of control

results from excessive aerodynamic moments exceeding thecontrol capability.

Vehicle tension loads due to the abrupt loss of thrust at

abort cutoff cause structural failure for actuator hardover

malfunctions which require abort between 70 and i00 seconds

flight time. Figure 4-30 shows tension loads for this

malfunction occurring at 80 seconds.

An S-IC actuator hardover occurring between 100 and 107

seconds flight time can cause loss of control due to the

excessive aerodynamic moment and the control gain switching

transient. Figures 4-31 and 4-32 show typical dynamics

and corresponding tension loads, respectively. No control

loss occurs and nostaging criteria are violated by anS-IC single actuator hardover between 108 and 160seconds.

During S-II flight, there is no loss of control for an actuatorhardover malfunction.

A single S-II actuator hardover inboard causes a large inboard

deflection of another control engine, as shown in Figure 4-33 •This exposes the base of the S-II vehicle to excessive

radiation from the engine plumes with the following possibleresults:

a. Collapse of thrust structure due to induced thermal stress

4-33

D5-15795-7

4.3.2 (Continued)

b . Loss of engine thrust and/or S-II/S-IVB separation

capability due to wiring harness damage.

S-II stage damage may occur within 15 to 25 seconds after

actuator hardover.

Any actuator hardover in the S-IVB stage results in tumbling.


Since there are no abort cues for early failures resulting

in holddown post or tower collision, it is assumed that thesefailures result in crew loss.

The manual abort cues and the automatic abort setting of

4 degrees per second provide positive abort lead times for

all S-IC actuator hardover failures in the high-q region

which require abort. Abort lead times are shown in Figure 4-34.

The figure shows that abort can be unsafe if an explosion

occurs at breakup, but is safe for all explosions occuring

one second after breakup. This results in the probability

of crew loss of 5 times per million flights. In most cases

the manual abort cues, attitude error (+5.0 degrees) and q-ball

AP (3.2 psid), occur before the automatic abort requirement

is reached.

S-IC actuator hardover between 108 and 120 seconds can result

in false automatic abort on overrate. Failures after 120

seconds do not require abort, and no false aborts are indicated.

While there is no loss of control for a single S-II engine

actuator hardover, abort cues for this malfunction are necessary

because of heating problems (Reference 6). These cues are

voice request and the Abort Request Light.

For actuator failures during S-IVB flight, the i0 degree/

second rate limit is the manual abort cue. All aborts are

safe.

4.3.4 Missfon Completion Capability

Any single actuator hardover in S-IC or S-II which does not

require abort for reasons discussed in Section 4.3.3 has

primary mission completion capability. Actuator hardoverin S-IVB results in mission loss.

4-34

D5-15795-7

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4.4.1 Malfunction

An actuator inoperative malfunction is any failure which

causes a single actuator to remain at or near null (±0.7

degree) regardless of control commands or external forces

placed on the actuator by the engine.


In S-IC and S-II flight, only minor dynamics due to thevehicle rotating to a trim attitude result from actuator

failures to null.

In S-IVB flight, an inoperative actuator causes attitude

divergence in the plane of the failed actuator. Figure 4-35

shows a typical case for a null pitch actuator. As the

failed actuator position approaches the limit for this

malfunction (±0.7 degree of null position), vehicle divergencebecomes more rapid.


Abort is not necessary for an actuator inoperative during S-IC

and S-II flights. First abort cue for S-IVB first and second

burns is attitude deviation. Second cue is one of the following:

a. Exceeding the ±45 degree yaw attitude limit

b. Exceeding the rate limits

c. Ground confirmation of attitude deviation (abort

request light) if received prior to the above.

Abort lead time is important for yaw actuator inoperative

because of the spacecraft platform yaw tumble limit of 90

degrees. Figure 4-36 shows sufficient lead time exists

for abort based upon the above cues.


Full mission completion capability is maintained for an

actuator inoperative during S-IC and S-II flight.

Abort is required during S-IVB first and second burns if the

malfunction occurs earlier than approximately 30 seconds

before nominal cutoff. If the malfunction occurs after this

time in second burn, full mission completion capability is

maintained; if it occurs before this time, TLI capability islost.

4-42

D5-15795-7

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D5-15795-7

4.5 SATURATED ERROR SIGNAL


Any failure that causes an erroneous 15.3 degree attitude error

signal is a saturated attitude error•


Figure 1-1 shows the times of malfunction that cause pad

fallback, liftoff interference, and tower collision• Since

the engines are not shut down in the abort sequence before

30 seconds, a saturated attitude error can turn the vehicle

and cause it to impact in the pad area.

Saturated attitude error in S-IC flight causes tumbling at all

flight times. Typical dynamics for this malfunction are shown

in Figure 4-37. Worst case launch vehicle tension loads that

result from a saturated control signal in the high-q region

are shown in Figure 4-38.

Saturation of the pitch or yaw error signal in the S-II stage

initially causes actuators hardover in the affected plane;

actuators hardover cause the vehicle attitude rate to increase

until the control equation (8 = ao_ + al_) becomes balanced.

This occurs when the actuators approach the null position, andthe attitude rate reaches a maximum value and remains at that

value. Shown below are the values used to calculate the

maximum rate for S-II saturated error signal for the first set

of S-II gains:

0 = (1.12)(15.3) + (1.89) _MAX (pitch and yaw)

_MAX (pitch and yaw) = -9.1 degrees/second

This maximum vehicle rate does not violate the EDS rate limit

of I0 degrees/second. However, the launch vehicle overrate

light may be activated due to the ±0.8 degree tolerance on the

±9.2 degrees/second limit setting. Attitude deviation of 20

degrees is exceeded within 5 seconds after time of malfunction•

Saturated roll error signal causes a large engine deflection

in both pitch and yaw planes• The vehicle roll rate reaches a

value which stabilizes the control equation. This value is

calculated below:

0 = (.25)(15.3) + (.2) _ (roll)

_MAX (roll) = -19 degrees/second

4-45

D5-15795-7

4.5.2 (Continued)

Saturated error signal in S-IVB produces the same effects asthose described for the S-II. Because of the difference incontrol gains ao and al, however, the EDS rate limit of 10degrees/second is exceeded for pitch and yaw malfunctions.

i

0 = (.81) (15.3) + (.97) _MAX (pitch and yaw)

¢MAx(pitch and yaw) = -12.8 degrees/second

0 = (i) (15.3) + (i) _MAX (roll)

%MAX (roll) = - 15.3 degrees�second

Figure 4-39 shows S-II and S-IVB pitch rate time histories

resulting from a saturated error signal malfunction.


Failures prior to 1.25 seconds result in the engines colliding

with the holddown posts. Crew loss is assumed for these cases,

since there are no valid cues for this situation. Failures

after 1.25 seconds result in aborts on angular rate cues. Abort

lead times for saturated error signal malfunctions during S-IC

flight are shown in Figure 4-40 . The figure shows a comparison

of actual lead time and required lead time for the most criti-

cal period of flight, from 30 seconds to 90 seconds. Lead times

during the remaining period of S-IC flight are much greater

than required. The figure indicates that if an explosion occurs

one second after breakup, the crew would be unsafe from explo-

sions following aborts occuring between 30 and 65 seconds.

In S-II, a launch vehicle overrate light may be used as a crew

abort cue; however, due to the balancing of the control equa-

tion, as described in Paragraph 4.5.2, a marginal condition

exists for the overrate light cue, necessitating reliance on

ground cues. Abort is by the launch escape system prior to

tower jettison and by the SPS after jettison; abort cues are

the abort request light and FDO limits.

Cues for saturated error signal in S-IVB are the same as those

in S-II. No automatic abort is provided for this malfunction

in S-II or S-IVB stages.


Saturated error signal in any stage of flight results in missionloss.

4-46

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D5-15795-7

4.6 SATURATED RATE SIGNAL


Any failure that causes a false attitude rate signal of i0

degrees/second to be sent to the flight control computer is

referred to as a saturated attitude rate signal.


Saturated rate signal causes loss of control at all times

during S-IC flight. Typical dynamics following this malfunc-

tion are shown in Figure 4-41.

Worst case launch vehicle tension loads that result from a

saturated rate signal in the high-q region are shown inFigure 4-42.

A saturated pitch or yaw rate signal in the S-II stage causes

the actuators in the affected plane to go hardover and the

attitude error increases until it reaches the limit of 15.3

degrees. The actuators then return the control engines to

a deflection of between 1 to 2 degrees, which results in

uncontrolled tumbling. A similar situation occurs for a

saturated roll rate signal. Figure 4-43 shows a vehicle

pitch rate history for a saturated rate signal during S-II

flight.

The effects of a saturated rate signal during S-IVB flight

are similar to those in S-II. Figure 4-43 shows an S-IVB

pitch rate history resulting from a saturated pitch rate

signal.


Failures prior to 1.25 seconds result in the engines colliding

with the holddown posts. Crew loss is assumed for these cases,since there are no valid cues for this situation. Failures

after 1.25 seconds result in aborts on angular rate cues.

EDS abort cues and logic for a saturated rate signal are

identical to those for a saturated attitude error (see

Paragraph 4.5.3).


Saturated rate signal in any stage of flight results inmission loss.

4-51

D5-1 5795-7

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4-54

D5-15795-7

4.7 LOSS OF INERTIAL ATTITUDE


An ST-124 inertial platform, located in the Instrument Unit,

is used as the inertial attitude reference during boost.

Platform outputs are monitored by the LVDC and tested for

reasonableness. On successive failures of the reasonableness

test, the LVDC issues a guidance failure discrete. This

discrete lights the LV GUID light on the spacecraft instrument

panel. It also turns o_ the LV overrate light prior toautomatic abort deactivation.

A backup system permits spacecraft takeover of guidance on

AS-507. Switchover to this backup system is performed

manually upon detection of the LV GUID light.

In this analysis it is assumed that the platform fails in

either the pitch, yaw, or roll plane at a constant angular

acceleration or at a constant drift rate. The gimbal reason-

ableness test is failed when the difference between successive

gimbal readings indicates a rate of change of gimbal angle

greater than 15 degrees per second. Figure 4-44 shows a

typical attitude error time history during the guidance

switchover sequence. Time t I depends on the angular

acceleration of the failed platform. Times between events

are based on the sequence given in Reference 7.


The vehicle attempts to follow a failed platform. Thus, a

high acceleration failure results in the vehicle accelerating

rapidly to follow it. Body rates in excess of abort rate

limits, which in this study are assumed to be equivalent to

loss of control, are thus reached in a very short time. In

all cases studied, aborts on rate preceded aborts on q-ball

AP_. Guidance switchover following platform failures in roll

can be accomplished throughout flight without loss of control.

Figures 4-45, 4-46, and 4-47 show the minimum platform

accelerations that are required to maintain control after

guidance switchover following positive pitch, negative pitch,and yaw platform failures, respectively.

Sensed attitude error at guidance switchover is not a

dependable abort cue since loss of control can result for

an error as small as 1.0 degree. Figures 4-48 and 4-49 show

the minimum attitude errors that can result at switchover

for cases that lose control. Errors are much larger during

the manual EDS mode because of the i0 degree per second abortrate limit.

4-55

D5-15795-7

4.7.2 (Continued)

_inimum accelerations for safe switchover during S-II flightcannot be defined because no dependable onboard EDS limits

are violated prior to activation of the LV GUID light. S-II

platform failure detection delays cause rapid buildup of

attitude errors to the saturation limit of 15.3 degrees.

This results in actuator trimming at vehicle rates just below

the i0 degree/second rate limit as described in Section 4.5.

Excessive attitude deviations at the time of LV GUID light

activation can cause violation of FDO limits subsequent to

S/C guidance switchover.

The maximum platform failure delay in S-IVB flight without

violating EDS abort limits is approximately 1.7 seconds.

Rate and error control gains in S-IVB flight do not cause

actuator trimming below the EDS rate abort limits as in S-II.

Very slow platform failures that do not violate the

guidance failure tests can result in off-nominal orbits.

Figure 4-50 shows maximum platform drift rates, with

resulting degradation of accelerometer data, that thevehicle can tolerate and still attain a 75-100 nautical

mile parking orbit. All non-immediate platform failures

with drift rates in excess of those defined in Figure

4-50 and with guidance failure delays exceeding 1.7seconds in S-IVB result in loss of vehicle.


As is shown by lead times in Figure 4-51, existing EDS auto-

matic and manual abort limits are adequate to ensure crew

safety during the transition from primary guidance to S/C

guidance. Slow platform failures in S-II may not violate

onboard EDS limits. These failures must rely on ground

monitoring of trajectory for violation of FDO limits andattitude deviations.


The launch vehicle can be guided to a safe parking orbit

(perigee altitude greater than 75 NMI) in the S/C backup

guidance mode and TLI can be achieved. Performance losses

due to manual guidance to insertion and translunar injection

are within the AS-507 reserves. An injection is achieved

requiring less than a 305 meter/second (i000 feet/second)

midcourse correction budget. These results are based on the

study of Reference 7.

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D5-15795-7

4.8 LOSS OF ATTITUDE COMMAND


Loss of attitude command includes all failures that result in

the attitude command remaining constant or becoming zero atthe time of failure.


Loss of roll attitude command can occur at all flight timeswithout causing mission loss.

Loss of yaw attitude command prior to approximately 350

seconds can result in excessive out-of-plane conditions such

that IGM fails to shut down the S-IVB engine at POI velocity.

This necessitates shutdown by ground command.

Constant pitch attitude command failures cause mission loss

during all flight stages except during the latter portions ofS-IVB first and second burns.

Failure of major loop pitch command to zero causes a one

degree/second pitch-up for boost-to-orbit. This malfunction

causes vehicle loss exceDt for late in S-IVB burns. If the

failure occurs prior to 90 seconds in S-IC flight, rapidvehicle tumbling results.

Failure of minor loop pitch command to zero causes a 12 degrees/

second pitch-up resulting in loss of control during S-IC flight.

Failures result in large attitude deviations during S-II and

S-IVB burns. Vehicle loss occurs for all failure times.


The abort cues for loss of yaw and pitch attitudecommand failures are:

a.

b.

Attitude deviation

Exceeding FDO limits, or ground confirmation of attitude

deviation (abort request light)

Failure of the major loop pitch command to zero prior to 90

seconds results in automatic abort based on 4 degrees/secondpitch rate. Manual abort cues are:

am

b.

Pitch attitude deviation

Exceeding FDO limits or ground confirmation of attitude

deviation (abort request light)

4-65

D5-15795-7

4.8.3 (Continued)

Failure of the minor loop pitch command to zero before 120seconds flight time results in automatic abort based on 4degrees/second pitch rate. After that time, manual abortcues are i0 degrees per second rate, and +5 degrees attitude

deviation during S-IC flight or +20 degrees attitude deviation

during S-II or S-IVB flight.

False abort or crew loss does not occur for loss of attitude

command malfunctions.


Mission completion capability exists for loss of roll command

at all flight times. Parking orbit insertion is achieved for

loss of yaw command failures after 350 seconds. Mission loss

occurs for a failure of pitch command to zero in boost-to-orbit.

For a constant pitch command occurring after 580 seconds of

flight time, POI is achieved, but TLI is not possible. In

S-IVB second burn, TLI is achieved for failure of pitch command

to zero after 320 seconds from reignition and for constant

pitch command after 270 seconds from reignition.

4-66

D5-15795-7

4.9 LOSS OF ATTITUDE ERRORSIGNAL


Any failure that causes a false attitude error signal of zerodegrees to the flight control computer is referred to as lossof attitude error signal.


For any malfunction time, loss of attitude error causesattitude divergence. Under pitch or yaw rate control only,the control system positions the thrust vector such that itcompensates for angular disturbances. This results in a vehiclerate dictated by the control law. During flight through theatmosphere, aerodynamic moments resulting from increasing angle-of-attack cause increasing rate. In the high-q region, un-controlled tumbling followed by structural failure occurs whenthe aerodynamic moment exceeds control authority. Duringupper stage flight, attitude divergence is slower and no struc-tural breakup occurs. Pitch and yaw failures cause loss of

vehicle except for the latter portion of S-IVB burn. No

vehicle loss occurs for roll failures. Figure 4-52 shows

typical dynamics for loss of pitch attitude error in S-IC.

Worst case loads for this case are shown in Figure 4-53.


During S-IC flight up to 120 seconds, safe aborts are provided

by the 4 degree per second overrate setting. Manual abortcues are:

a.

b.

C.

5 degrees attitude error

3.2 psi q-ball pressure (between 50 and 120 seconds)

i0 degree per second overrate (after 120 seconds)

Abort cues for failures during S-II and S-IVB flight are:

a.

b.+20 degrees attitude deviation

FDO limits or ground confirmation of attitude deviation

(abort request light)

False abort or crew loss is not expected for loss of attitude

error signal malfunctions.


Loss of attitude error (pitch or yaw) during S-IC, S-II, or

early S-IVB flight results in loss of mission. Parking orbit

insertion is attained for loss of pitch or yaw error after

620 seconds. In S-IVB 2nd burn, full mission completion capabi-

lity exists for loss of pitch or yaw error signal after approxi-

mately 310 seconds from reignition.

4-67

D5-15795-7

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:'.i0 LOSS OF ATTITUDE RATE SIGNAL


Any failure that causes a false vehicle rate indication of

zero degrees per second to the flight control computer is

referred to as a loss of attitude rate signal.


Except for roll failures during S-IVB flight, the loss of

attitude rate signal causes oscillatory divergence of vehicle

attitude due to the absence of rate damping. During S-IC

flight in the high-q region, this divergence is aggravated

by the buildup of aerodynamic moments, eventually resulting in

structural breakup. Dynamic variables for a loss of rate

signal at 65 seconds are shown in Figure 4-54 . Loads that

result from this malfunction are shown in Figure 4-55 • Figure

4-56 shows dynamics for loss of pitch rate signal during S-II

flight.

Loss of roll rate signal in S-IVB produces only small attitude

oscillations which do not diverge. This results from the

different method of roll axis control used during S-IVB flight,

namely, the Auxiliary Propulsion System (APS). The roll

attitude of the vehicle is maintained within the 1 degree

deadband of the APS control system as shown in Figure 4-57.


During S-IC flight, safe automatic aborts are provided by the

4 degree per second overrate setting and rate indicator. After

automatic abort is disabled, manual abort cues are i0 degrees

per second overrate and 5 degrees attitude error.

Manual abort cues for failures during S-II and S-IVB flight

are the i0 degree per second overrate light and spacecraft rateindicator.

False aborts or crew loss do not occur for loss of rate

signal malfunctions.


Loss of attitude rate during S-IC and S-II stage flight results

in vehicle and mission loss. For loss of pitch rate signal

as early as 35 seconds before S-IVB engine cutoff, and loss of

yaw rate signal as early as i00 seconds before S-IVB engine

cutoff, insertion or injection may be achieved with residual

rates less than i0 degrees per second.

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4.11 ACCELEROMETER MALFUNCTIONS


Failures considered are those which result in a zero acceler-

ometer reading on a single axis with and without thrust

misalignment and the effects of manual S-IVB shutdown and

navigator update.

The accelerometer output is tested for reasonableness in

each major cycle. When the zero test is enabled (during

each stage mainstage burn) an accelerometer output of less

than 10.051 meters/second is accepted when the vehicle

longitudinal axis is within A@ degrees of normal to that accel-

erometer input axis. A0 is 2 degrees except during S-IC and

S-II burn with a premature outboard engine shutdown, when it

becomes 6 degrees. Backup data is substituted by the acceler-

ometer processing logic when the accelerometer reading isrejected.

4.11.2 Malfunction Without Thrust Misalignment

Into-orbit X-accelerometer failures for a nominal vehicle

result in perigee altitudes greater than 75 nautical miles.

For +3o and -30 vehicles (30 acceleration profile), X-acceler-

ometer failures result in perigee altitudes greater than 75

nautical miles after 135 and 30 seconds, respectively. Theworst case is a -3o X-accelerometer failure at 2.0 seconds

which results in a perigee altitude of 53.5 nautical miles.

Apogee/perigee altitudes for X-accelerometer failures are

presented in Figure 4-58.

Into-orbit Z-accelerometer failures result in elliptical

orbits with perigees near 100 nautical miles. Figure 4-59

shows apogee and perigee altitudes for nominal, +3o, and-3o Z-accelerometer failures.

Into-orbit Y-accelerometer failures result in nearly circular

i00 nautical mile orbits. Yaw steering angle history for aY-axis accelerometer failure at 2 seconds after first motion

is presented in Figure 4-60 for boost to insertionand in

Figure 4-61 for boost to TLI. For this failure, the earth

orbit has a -0.1443 degree error in inclination and a -0.2177

degree error in descending node. The Y-axis velocity errorat TLI is 20 meters/second.

The primary analysis for out-of-orbit accelerometer failures

is for a September 13 (1969) launch date, a 78.051 degree

launch azimuth, and a first opportunity. The velocity error

in the failed axis (from desired cutoff hypersurface) at TLI

for failure times of an X, Y, or Z-accelerometer is presented

for the primary analysis in Figures 4-62 through 4-64. This

error, rather than the perturbation from the nominal cutoff

4-75

D5-15795-7

4. ii. 2 (Continued)

velocity is presented because following an accelerometerfailure, the hypersurface cutoff conditions differ from the

nominal. The velocity components normal to the failed axis

have no error from the targeted cutoff hypersurface other

than normal accelerometer output error. Steering angle

envelopes for all failure times and nominal histories for

pitch angle and yaw angle are presented for the primary

analysis in Figures 4-65 and 4-66, respectively.

The velocity error in the failed axis (see preceding paragraph)

for first opportunity and across the launch window is shown in

Figures 4-67 and 4-68 for X- and Z-accelerometer failures.

Both maximum error and error as a function of failure time

are obtainable from the figures for different days and azimuths.

Velocity error for Y-accelerometer failures varies relative

to how long IGM rides the 2 ° deadband and has a maximum

predicted error of the order of 50 meters per second.

4.11.3 Malfunction With Thrust Misalignment

A thrust misalignment with no accelerometer failure has

negligible effect on the flight because a calculation,

Steering Misalignment Correction (SMC), is provided to

correct the steering commands for the null input deflection

error.

The insertion or injection error occurring with a pitch thrust

misalignment concurrent with a pitch plane accelerometer

failure is approximately equal to (±5 meters/second) the error

for the accelerometer failure with no misalignment. The

misalignment causes a slightly revised steering history

resulting in a different boost duration compared with a

similar accelerometer failure with no misalignment. This

boost duration change ranges up to 2 seconds at POI, 3_

thrust misalignment in all three stages for into-orbit

accelerometer failures, and 0.5 seconds at TLI, for S-IVBsecoad burn failures.

The reason for these effects is that, while the misalignment

causes the control system to set up a delta between the

commanded (X) and actual (8) attitudes with which to zero out

the misalignment, i.e.,

B = B O + a(@-X) + b_

where: a = altitude gain

= rate gain

= vehicle body rate

B = engine deflection

Bo = thrust misalignment,

4-76

D5-15795-7

4.11.3 (Continued)

the navigator sees the backup acceleration resolved throughthe actual attitude angles. The control system (with orwithout a misalignment) maintains the thrust through thevehicle CG. The angular error between this thrust accelerationand axial acceleration and also the magnitude error betweenthe backup and actual acceleration is the same with or withouta misalignment.

The insertion error with a yaw thrust misalignment concurrentwith a Y-axis accelerometer failure depends on how long the 2°accelerometer output zero test limit is followed. A ±3athrust misalignment causes up to 65 m/second additional errorin Y at POI compared with a similar accelerometer failure withno misalignment and up to 0.5 second extension of the S-IVBfirst burn duration. The injection error with the samevehicle condition results in an additional error in Y at TLIof up to 5 m/second and up to 0.5 second extension to theS-IVB second burn. If a navigator update is used to correctan into-orbit Y error, then the into-and out-of-orbit Y errorsare not additive. Note that the 2 degree zero test deadbandwill also affect certain day/azimuth S-IVB second burn _ andZ-accelerometer failures when the vehicle major axis lieswithin 2 degrees of the failed axis during second burn.

4.11.4 Manual S-IVB Shutdown

A manual shutdown on inertial velocity can reduce overspeedat POI and TLI. The manual cutoff at POI will only reducethe overspeed produced by an early Z-axis accelerometerfailure. X-axis accelerometer failures have more of a flightpath angle error than an overspeed and a manual cutoff doesnot reduce their POI error. Similarly, Y-axis failures havean orbit inclination error and little overspeed. A navigatorupdate is required in EPO to correct disagreement between theactual and navigator state vectors.

At TLI with an LVDC cutoff, the overspeed is dependent on theS-IVB reignition angle (position vector in the X/Z plane) andthe failure time (i.e., the component of the FOM-FOMCerror inthe failed axis). The overspeed reduction achieved by amanual shutdown on inertial velocity is dependent on thedirection of the total acceleration immediately prior to TLI.The inertial velocity specified for the manual shutdown isobtained from a realtime S-IVB second burn simulation fromthe achieved EPO.


Abort may be required for X-axis accelerometer failures priorto 30 seconds after first motion on a -3a vehicle and 135

4-77

D5-15795-7

4. ii. 5 (Continued)

seconds after first motion on a +3_ vehicle. The abort cuesare FDO limits and abort request light. A Mode IV abort isrequired.


_ission completion capability for accelerometer failuresis shown in Figure 1-9; mission continuance will be a real

time decision. All out-of-orbit accelerometer failures

result in perturbed TLI injection requiring midcourse

correction and/or manual cutoff.

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4.12 PU SYSTEMMALFUNCTIONS

.12.1 Malfunction Description

The S-II and S-IVB Propellant Utilization (PU) System mal-functions are those in which the PU valve fails to followthe flight plan. No malfunctions are considered to occur be-fore PU unlock. All valves fail simultaneously to somefixed position for remainder of flight.


The primary effects of S-II PU malfunctions are trajectorydeviations and perturbed S-IVB first burn times. Trajectoryfor PU failures to null and low stop exceed the nominal atS-IVB insertion into parking orbit. S-IVB PU malfunctionscause less deviation from the nominal trajectory than S-IIPU malfunctions.


Aborts are not required for PU malfunctions, and no falseaborts are initiated.


Parking orbit insertion can be achieved for all PU mal-functions in boost-to-orbit. Translunar injection (TLI) can-not be achieved for PU malfunctions to null position or lowstop occurring in the early portion of S-II flight. For afirst opportunity reignition, PU malfunctions to the lowstop in the later portion of the S-IVB first burn and all ofthe S-IVB second burn achieve TLI. For a second opportunityreignition, PU malfunctions to the low stop in S-IVB firstburn and the early portion of S-IVB second burn will notachieve TLI. Failure to achieve TLI is due to propellantdepletion. Table 4-I presents AS-505 preflight pre-dictions applicable to AS-507. Failure times indicated arethe earliest for which TLI may be achieved. "N/A" in thetable indicates the PU valve is in its normal operating posi-tion. "All" in the table indicates vehicles with failuresto the corresponding valve position have TLI capability forall failure times.

4-90

D5-15795-7

TABLE 4-I. TLI CAPABILITY FOR PU MALFUNCTIONS

STAGE

S-II

S-IVB First OpportunityFirst BurnSecond Burn

S-IVB Second OpportunityFirst BurnSecond Burn

FAILURE TONULL

POSITION

340 Sec

NIAAll

N/AN/A

FAILURE TOLOWSTOP

475 Sec

370 SecAll

No TLI14,610 Sec

4-91

D5-15795-7

4.13 LOSS OF ONE APS MODULE

_.13.1 Malfunction Description

Any nozzle firing failure in the APS system which causes lack

of APS response to firing commands in the S-IVB stage is de-

fined as APS System Failure. Other failures upstream of the

control computer output are not considered here.

±Joss of one APS module is defined as the failure of all

thrusters in one APS module to fire when commanded to do so.

APS ullaging failures are defined as malfunctions resulting in

failure to turn on an APS ullage thruster. Only single failuresare considered.


Loss of one APS module in S-IVB first or second burn will not

result in abort conditions; pitch and yaw control are main-

tained by the main engine and the remaining module provides

sufficient roll torque to correct for small perturbations.

Loss of one APS module during coast flight is cause for abort,

since control about the pitch axis exists in only one direction.

This malfunction during S-IVB coast will also cause a gradual

roll/yaw attitude divergence.

Loss of one APS ullage motor does not cause loss of control.


Loss of one APS module during parking orbit produces

loss of control in the pitch axis. Since attitude divergence

is slow, APS failure may not be detected immediately. The

abort cues are +20 degrees attitude deviation and ground

requested abort? All aborts are safe aborts.

False abort or crew loss does not occur for loss of one APS

module.


While loss of a single APS module during S-IVB first burn

allows a nominal POI, an abort during parking orbit is

required. Abort is required if this malfunction occurs

during parking orbit. Loss of one APS module during S-IVB

second burn does not prevent TLI; subseqUent to TLI, however,

no pitch control is available. This lack of control results

in uncorrected vehicle body rates.

4-92

D5-15795-7

4.14 LOSS OF BOTH APS MODULES


"Loss" is defined as in Paragraph 4.13.1, except both APSmodules are affected.


Loss of both APS modules in S-IVB powered flight produces lossof roll control.

Loss of both APS modules during S-IVB parking orbit coast

causes total loss of attitude control and requires an abort.

Although TLI is attained for loss of both APS modules during

S-IVB second burn, rates continue uncorrected, presenting

difficulties for transposition, docking, and extraction (TD&E).


Loss of both APS modules during S-IVB powered flight results

in loss of roll control. Excessive roll attitude is used as

a cue for manual abort.

For both APS modules out during coast flight, the S-IVB is

uncontrollable, but divergence is slow. Since body rates are

small, this malfunction may not be detected immediately. Abort

should be initiated before platform gimbal limits are exceeded.

Abort cues for loss of both APS modules are excessive attitude

deviation and abort request light.

False abort or crew loss is not expected for loss of both APSmodules.


Loss of both APS modules during S-IVB first burn or parking

orbit coast results in mission loss. During S-IVB second burn,this malfunction will not prevent TLI; but the S-IVB will

be uncontrollable after TLI, and capability to perform theTD&E is questionable.

4-93

D5-15795-7

4.1 r SEQUENCINGAND STAGING MALFUNCTIONS

_.15.1 Malfunction Description

A sequencing malfunction is defined as any malfunction thatcauses an erroneous sequence or no sequence to occur, or

a sequence that occurs at the wrong time. Sequencing failuresstudied are:

Do

Early gain switching

Late (or no) gain switching.

A staging malfunction is defined as any malfunction that

causes premature staging, lack of staging, or complications

during staging. Those studied include:

a. Lack of retrorocket firing

b. Lack of (or partial) S-IC first plane separation

c. Lack of (or partial) S-II second plane separation

d. Lack of (or partial) S-II/S-IVB plane separation

e. Failure to jettison the Launch Escape Tower (LET)


An early control system gain change (pr±or to 85 seconds)results in vehicle loss because the vehicle is unstable in

the max-q region with low gains.

A late gain change does not require abort even if no gain

change occurs during S-IC flight. The vehicle becomes un-

stable due to slosh, but more than 50 seconds of instability

are required to cause loss of control.

Failure to achieve S-IC/S-II staging requires abort since

parking orbit insertion cannot be attained by staging

directly to the S-IVB.

One retrorocket out at S-IC/S-II or S-II/S-IVB staging does

not require abort. Failure to issue firing commands to the

S-IC or S-II retrorockets produces delayed stage separation

and a possible period of local deformation where the stages

remain in contact after they have been severed. Thrust

tailoff forces of the engines tend to keep the stages in

contact; precise structural consequences are not known.

A sequencing failure affecting time of gain change does not

cause an abort during S-II flight.

4-94

D5-15795-7

4.15.2 (Continued)

Sequencing failures or separation device malfunctions whichprevent jettison of the S-IC/S-II interstage cause the thermalenvironment limits in the S-II boattail area to be exceeded.Engine shutdown and abort are required prior to generation ofexcessive temperatures. Due to the short time between nominalinterstage jettison and exceeding termperature limits, themission rule presented in Reference 6 (abort prior toTB3 + 52 seconds) is recommended.

Failure to jettison the launch escape tower introduces nostability problem during S-II and S-IVB flight. The launchescape tower and boost protective cover prevent spacecraft/lunar module transposition and docking. Reentry is notpossible since the drogue and main parachutes cannot bedeployed. Procedures are required to allow the crew to removethe LET and boost protective cover during earth parking orbit.

Failure to achieve S-II/S-IVB staging due to a failure toissue the sequence command requires abort.

An early gain change or absence of one during S-IVB secondburn does not require abort.


Table 4-IIcontains the abort logic and limits for all sequencingand staging malfunctions considered in this analysis. Allabort lead times are adequate for crew safety.

False abort or crew loss does not occur for sequencing and

staging malfunctions.


Mission completion capability for sequencing and stagingmalfunctions is given in Column 4 of Table 4-I.

4-95

D5-I'5795-7

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4-96

D5-15795-7

4.16 S-II/S-IVB EARLY STAGING

4.16.1 Description

.

Manual early staging to alternate timebase 4 sequence can

be initiated between thrust OK arm, 6.7 seconds after

timebase 3, and propellant depletion cutoff arm, 334.2seconds after timebase 3.

4.16.2 Limits

The largest body rate from which early staging can be initiated

and not fail the fine gimbal resolver test is approximately 2

degrees/second. The rate from which successful S-IVB recovery

can be anticipated after early staging should not exceed3 degrees/second.


S-II/S-IVB early staging capability is shown in Figure 1-8.

S-IVB burn duration as a function of the time of early stagingis presented in Figure 4-69.

4.16.4 African Continent Impacts

Table 4-III details the intervals of S-II/S-IVB early stagingtimes for a nominal vehicle which result in African continent

impact. These intervals are projected from AS-507 data andAS-505 results.

TABLE 4-111 S-II/S-IVB EARLY STAGING CONDITIONSFOR AFRICAN CONTINENT IMPACTS

I

LAND IMPACTAREA

72 ° AZIMUTHWEST COAST(CANARYISLANDS)

EAST COAST

90° AZIMUTHWEST COAST(CAPE VERDEISLANDS)

EAST COASTI

108 ° AZIMUTHWEST COAST

EAST COAST

STAGINGTIME t

SECONDS

330

375*

330

375*

365

375*

GEOCENTRICRADIUSMETERS

6545526.

6555080.

6546123.

6555673.

6553990.

6555628.

INERTIALVELOCITYM/SEC

3995.

4516.

4014.

4536.

4392.

4516.

INERTIAL FLIGHTPATH ANGLEDEGREES

3.983

1.936

3.973

1.918

2.303

1.92Z

tNOTE: These times are qiven to the nearest 5 seconds.*Achieves parking orbit.

4-97

D5-15795-7

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4-98

D5-15795-7

SECTION 5

COMMUNICATIONS ANALYSIS FOR ABORT AND ALTERNATE MISSIONS

5.1 COMMUNICATIONS ANALYSIS BACKGROUND

The objective of this study is to provide an assessment of the

tracking and communications capability, for a defined sur-

veillance network (Table 5-I), for abort and alternate missions

within the capability of the AS-507 vehicle. This study

complements analyses presented in Boeing Document D5-15697-507,

"Tracking and Telemetry Design Analysis for AS-507 (G Mission),"

dated June 20, 1969. Abort and alternate mission trajectories,

resulting from vehicle system or subsystem malfunctions,

result in deviations from the nominal communications capability

such that additional data are required for contingency planning.

In the referenced document, an extensive analysis of the

tracking and communications capability is presented for the

nominal AS-507 G Mission based upon the AS-507 Launch Vehicle

Operational Flight Trajectory. A summary of the nominal

0-degree elevation angle surveillance for the Earth Parking

Orbit is given in Table 5-II for the 78.051-degree launchazimuth on September 13, 1969.

The nominal tracking and communications analysis for the

launch and parking-orbit flight phases is based upon the

AS-506 Launch Vehicle Operational Flight Trajectory. The

AS-507 Launch Vehicle Operational Trajectory during those

flight phases does not show any appreciable differences, in

regards to geometrical surveillance, from AS-506 trajectory

data. Computer runs were made using the AS-507 Operational

Trajectory to verify the applicability of the AS-506

surveillance data. The alternate mission analysis is basedon the AS-507 launch vehicle.

An abort mission, for purposes of this analysis, is a mission

that results when any vehicle system or subsystem failure

prevents the launch vehicle from achieving insertion into a

parking orbit; therefore, the communications analysis of the

abort missions is restricted to the launch and boost-to-

parking-orbit phase of flight.

An alternate mission is, for purposes of this analysis, any

mission resulting from a system or subsystem failure that

permits insertion of the launch vehicle into an earth orbit.

Alternate missions are analyzed during both the boost and

earth parking orbit flight phases to develop surveillance

data for use in an overall assessment of the tracking and

communications capability for the AS-507 G Mission.

5-1

D5-15795-7

5.1 (Continued)

A detailed tracking and communications analysis for all ofthe possible abort and alternate missions to the detailprovided by the referenced AS-507 G Mission nominalsurveillance analysis is impractical. To satisfy the needfor displaying meaningful data for the various possibleabort and alternate missions, analyses are conducted toidentify fundamental tracking properties applicable to thevarious flight phases.

A special surveillance analysis is included in this sectionthat provides surveillance for both the instance of no S-IVBstage restart ignition (for variable azimuth across the launchwindow) and for the instance of an early stage cutoff forSeptember 13, 1969, second TLI opportunity and 78.051-degreelaunch azimuth.

No allowances are made in this study for the effects thatresult from a degradation in antenna look-angles. Anydegradation in antenna look-angles may result in a change inthe capability to communicate with the launch vehicle.

5.2 ANALYTICAL PROCEDURESAND STUDY LIMITATIONS

During the launch and boost-to-parking-orbit flight phase,continuous tracking and communications surveillance isavailable for all defined alternate mission trajectories.Because of the diversity of the altitude/surface rangeprofiles for abort trajectories, each abort condition must beconsidered as an individual communication problem; therefore,specific abort cases are not presented in this analysis.

Any trajectory with an altitude/surface range profile abovethe surface of surveillance defined by the minimum altitude/surface range contours shown in Figure 5-1 is undersurveillance continuously by one or more Eastern Test Rangestations. This surface of surveillance is generated byjoining the points of intersection of the adjacent trackingand communications station visibility cones of the MILA,Grand Bahama, Grand Turk, Antigua, and Bermuda groundstations. Limited surveillance is available below thissurface, but continuous surveillance is unavailable.

The AS-507 G Mission nominal altitude/surface range profileand the minimum altitude/surface range profile for analternate mission, resulting from an S-IC stage dual-engineshutdown at 90 seconds after liftoff, are presented inFigure 5-1 to indicate nominal and typical worst-casemalfunction flight profiles that achieve a parking orbit.

5-2

D5-15795-7

5.2 (Continued)

The data of Figure 5-1 are valid for a launch azimuthvariation from 72 degrees to 108 degrees east of north. Thenominal and S-IC dual-engine-out altitude/surface rangeprofiles shown are valid across the launch azimuth spread.

Use of the surveillance contours defined in Figure 5-1 doesnot identify specific acquisition and loss times or parametersfor each possible abort or alternate mission. Analysis hasshown, however, that coverage for all alternate missionsresembles predictions made for a nominal flight during theboost flight interval.

The earth parking orbits that result from alternate missionsare classified into the following categories:

a. Alternate missions that achieve a near-nominal earth

parking orbit.

b. Alternate missions that achieve an off-nominal

elliptical parking orbit.

The alternate mission trajectories that achieve a near-

nominal earth parking orbit result from propulsion system

or subsystem failures that reduce the vehicle's overall

performance capability. These trajectories are discussed in

detail in Paragraph 5.3.

Alternate missions that achieve an elliptical earth parkingorbit are the result of accelerometer malfunctions. An

analysis of the alternate missions, resulting from acceler-

ometer malfunctions, is presented in Paragraph 5.4.

The surveillance analysis applicable to the particularinstance where no S-IVB second burn occurs consists of extend-

ing the nominal earth parking orbits for the AS-507 G Mission.

Additional analyses are provided to define surveillance for

September 13, 1969, 78.051-degree launch azimuth, second

opportunity for the special instance of an S-IVB stage early

cutoff. This analysis is discussed in Paragraph 5.5.

A study of station acquisition azimuths is included to augment

the station acquisition aid systems that are limited to a 20-

degree-wide antenna pattern. This study includes a presenta-tion of all acquisition azimuths that deviate more than i0

degrees from the corresponding nominal azimuth during thecrltical first orbit after insertion.

5-3

D5-15795-7

5.3 SURVEILLANCEFOR NONACCELEROMETERMALFUNCTIONS

Alternate missions resulting from nonaccelerometer malfunctionsare defined in Table 5-III and are classified into thefollowing types:

a. Single and dual S-IC and S-II stage engine malfunctions.

b, Malfunctions resulting in early staging from the S-II

to the S-IVB stage.

The earth parking orbits resulting from the defined failures

lie in the plane of, and closely approximate, the nominal

earth parking orbit as defined by the AS-507 G Mission

operational trajectory. The groundtracks of these orbits

are given in Figures 5-2, 5-3, and 5-4. Each of the

nonaccelerometer failures, however, achieves earth parking

orbit insertion at a later flight time than does the nominal.

The difference in station acquisition and loss times, obtained

by comparing a delayed-nominal orbit with the corresponding

nominal parking orbit, is referred to as delta time. The

delta time that is related to the time that a nonaccelerometer

malfunction occurs is approximately constant from station to

station for any orbit or for any launch azimuth.

The delta time for a particular time of nonaccelerometer

malfunction is computed by recording the time and longitude

after the malfunctioning vehicle has achieved orbital

insertion and by differencing this time with the time at

which the corresponding nominal vehicle reaches the same

longitude in the nominal parking orbit.

The delta times computed for nonaccelerometer failures that

occur between liftoff and nominal S-II cutoff are provided

in Figure 5-5. The acquisition and loss times for a delayed

parking orbit are determined by addinq the delta time from

Figure 5-5 to the nominal acquisition and loss times of

Figures 5-6, 5-7, and 5-8.

The endpoints (designated by solid lines) of the delta time

curves for the nonaccelerometer failures shown in Figure 5-5

represent the extent of the trajectory data available for

analysis. The analysis of the engine-out and early staging

malfunctions begins with the earliest malfunctioning vehicle

that achieves parking orbit insertion.

The dashed lines on the delta time curves in Figure 5-5

represent an extrapolation from the last actual trajectory

data available to the known terminal condition.

5-4

D5-15795-7

5.3 (Continued)

In this analysis, the delta time is computed for the first

orbit and is a good approximation of the delta times for thesecond and third orbits.

No significant acquisition azimuth deviations from the

nominal occur for the nonaccelerometer malfunctions.

5.4 SURVEILLANCE FOR ALTERNATE MISSIONS RESULTING FROMACCELEROMETER FAILURES

The specific measurements normally supplied to the guidance

and control system by an accelerometer are replaced by

guidance-computed estimates when an accelerometer malfunction

occurs. The errors in these guidance estimates, in general,

cause inplane overspeed and flight-path-angle errors at

orbital insertion. The resultant orbits are characterized

by apogees varying between 185 and 1815 kilometers and byperigees varying between 99 and 187 kilometers.

The apogee and perigee altitudes resulting from X- and

Z-accelerometer malfunctions are described in Figures 4-58

and 4-59. These altitudes are a function of accelerometer

malfunction time during the boost phase for both a nominal

and ±3a performance vehicle. The property of most

importance to the tracking and communications analysis

is that malfunctions during boost result in earth parkingorbits that have related orbital properties.

The Z-accelerometer malfunctions have related parking orbits

since the resultant perigees are approximately a single-

valued function of apogee. The X-accelerometer malfunctions

do not have this property, and the surveillance data arepresented as a function of X-accelerometer time ofmalfunction

Station acquisition and loss flight time histories for

X-accelerometer failures occurring on a nominally performingvehicle are presented in Figures 5-9 through 5-12. These

data are presented as a function of failure time after

liftoff. The data of these figures approximate the

acquisition and loss times for the ±3a performance vehicle

for malfunctions occurring after 60 seconds. An early

X-accelerometer malfunction occurring on a -3a performancevehicle results in an abort situation.

5-5

D5-15795-7

5.4 (Continued)

A 10-degree deviation in acquisition azimuth from the nominalsurveillance occurs only for Honeysuckle. The acquisitionazimuth as a function of time of X-accelerometer malfunctionis described in Figure 5-13 for this surveillance station.

Station acquisition and loss flight time histories arepresented for Z-accelerometer (both ±3a and nominallyperforming vehicles) failures as a function of orbitalapogee altitude. (See Figures 5-14 through 5-17.) Thesedata are related to an accelerometer malfunction as follows:

a . The apogee altitude is determined for a malfunction at a

specific time during launch by using the data of

Figure 4-59.

b. The station acquisition and loss time is determined

as a function of the apogee altitude by using the data of

Figures 5-14 through 5-17.

A 10-degree deviation in acquisition azimuth from the nominal

surveillance occurs for the Tananarive, Carnarvon, Honeysuckle,

and Goldstone tracking stations. The acquisition azimuths for

these stations is provided as a function of apogee in

Figure 5-18.

5.5 S-IVB SECOND-BURN EARLY CUTOFF SURVEILLANCE ANALYSIS

The analysis for the S-IVB second-burn early cutoff is divided

into two separate studies. The first study concerns the

instance of no S-IVB stage restart ignition and is the most

likely malfunction of this kind. The surveillance analysis

for the no-restart stage ignition is provided across the

launch window for the launch-azimuth variation from 72 degrees

to 108 degrees. The data provided in Figures 5-7, 5-8, and in

Figures 5-19 through 5-21 define the surveillance for the in-

stance of no S-IVB stage reignition and is based upon the con-

tinuation of the earth parking orbits for the AS-507 G Mission.

The second analysis describes the type of tracking and com-

munications surveillance expected for a typical launch day,

azimuth, and TLI opportunity for an S-IVB early cutoff

occurring at any instant during the second burn. This par-

ticular study was completed for the September 13, 1969,

second TLI opportunity and for a 78.051-degree launch azimuth.

(See Figure 5-22.) These data are shown only to 19,000

seconds because of the difference in the orbital period for

orbits resulting from the different failure times. The

acquisition azimuths for an early S-IVB second burn cutoff

for the Canary Island and Madrid stations deviate by more

than i0 degrees from the nominal second-burn and post-TLI

5-6

D5-15795-7

5.5 (Continued)

surveillance. The acquisition azimuths for these twostations as a function of time of malfunction during thesecond burn are given in Figure 5-23. Acquisition azimuthsfor the special instance where no S-IVB stage second burnoccurs is provided in Table 5-II.

The Y-accelerometer malfunctions result in out-of-planeerrors at insertion. A limited analysis of these malfunctionsis represented by the station acquisition and loss historyfor a typical worst case. (See Table 5-IV.) No acquisitionazimuth deviations exceed 10 degrees from the nominalsurveillance.

5-7

D5-15795-7

TABLE 5-I. SURVEILLANCENETWORKUSEDFOR AS-507 G MISSIONCOMMUNICATIONSANALYSIS FOR ABORTAND ALTERNATEMISSIONS

STATION

MERRITT ISLAND

GRANDBAHAMA

BERMUDA

ANTIGUA

INSERTION SHIP

CANARYISLAND

ASCENSION

MADRID

TANANARIVE

CARNARVON

GUAM

HONEYSUCKLE

HAWAII

GOLDSTONE

GUAYMAS

CORPUSCHRISTI

SYSTEM

CCS,TMCCS,TMCCS,TMCCS,TMCCS,TMCCS,TMCCS,TMCCS

TM

CCS,TMCCS,TMCCS

CCS,TMCCS

CCS,TMCCS,TM

SYMBOL

MIL

GBM

BDA

ANT

INS

CYI

ASC

MAD

TAN

CRO

GWM

HSK

HAW

GDS

GYM

TEX

GEODETICLATITUDE

(DEGREES)

28.508272

26.632857

32.351286

17.016916

25.00000

27.764536

- 7.955056

40.455358

-19.000797

-24.907592

13.309244

-35.597222

22.124897

35.341694

27.963206

27.653750

LONGITUDEEAST

(DEGREES)

- 80.693417

- 78.237664

- 64.658181

- 61.752849

- 49.000000

- 15.634814

- 14.327578

- 4.167394

47.315053

113.724247

144.734414

148.979167

-159.664989

-116.873289

-110.720850

- 97.378469

HEIGHTABOVE

ELLIPSOID(METERS)

i0.00

5.00

21.00

43.00

0.00

173.00

562.00

825.00

1322.30

58.00

127.00

1097.00

1150.00

965.00

19.00

10.00

NOTE: ALL GEODETIC DATA ARE REFERENCEDTO THE FISCHERELLIPSOID.

CCS -- COMMANDAND COMMUNICATIONSSYSTEM

TM -- TELEMETRY

5-8

D5-15795-7

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Table 5-III° ALTERNATEMISSIONS ACHIEVING A NOMINALPARKING ORBIT

SYMBOLSUSED ONALL DATA PLOTS

S-IC 1 EO

S-IC 2 EO

S-IC 1+5 EO

S-IC 2+4 EO

S-II 1 EO

S-II 1+5 EO

S-II 1+4 EO

S-II 2+4 EO

S-II/S-IVB ES

TYPE OF FAILURE

Single-engine-out (Engine No. i)failure in the S-IC stage.

Single-engine-out (Engine No. 2)failure in the S-IC stage.

Dual-engine-out (Engines No. 1 andNo. 5) failure in the S-IC stage.

Dual-engine-out (Engines No. 2 andNo. 4) failure in the S-IC stage.

Single-engine-out (Engine No. l)failure in the S-II stage.

Dual-engine-out (Engines No. 1 andNo. 5) failure in the S-II stage.

Dual-engine-out (Engines No. 1 andNo. 4) failure in the S-II stage.

Dual engines out (Engines No. 2 andNo. 4) on the S-II stage.

Early staging of the S-II to theS-IVB stage.

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Documents

· DOCUMENT NO. D5-15795-7 TITLE SATURN V AS-507 LAUNCH VEHICLE OPERATIONAL ABORT AND MALFUNCTIONED FLIGHT ANALYSIS MODEL NO. SATURN V …