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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
S-IC Sequential Engine-Out Capability - EnglnesNo. 2 and No. 4
S-IC Sequential Engine-Out Capability - EnglnesNo. 3 and No. 5
S-IC Sequential Engine-Out Capability - EnglnesNo. 2 and No. 5
Mission Completion Capability for S-II No. 1
& 2 Engine Shutdown
Mission Completion Capability for S-II No.
1 & 4 Engine Shutdown
Mission Completion Capability for S-II No.
1 & 5 Engine Shutdown
Mission Completion Capability for S-II No.
2 & 3 Engine Shutdown
Mission Completion Capability for S-II No.
2 & 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
ILLUSTRATIONS (CONTINUED)
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)
Time History of Dynamic Variables for Loss of 4-71
Attitude Rate (S-IC)
Tension Load at Sta. 1564 (Loss of Attitude Rate) 4-72
Time History of Dynamic Variables for Loss of 4-73
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
ILLUSTRATIONS (CONTINUED)
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
78.051 ° Launch Azimuth
0-Degree Acquisition and Loss History for Fourth
Orbit for Z-Accelerometer Malfunctions for
78.051 ° Launch Azimuth
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
•
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
I
FIGURE 2-i
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1'0 1"I 12 1'3
K
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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
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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'-]
I
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I ATTITUDERATE
I
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C
J VEHICLE
I ANGULAR RATE
---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
t
,qI'---ATTITUDEERROR
_1 COMBINE & FILTER INPUTS
•_[ TO GETCONTROL _ ANGLES i IIJ
BRc' BPc, & BYc
TVC FLUID 1POWER SUBSYSTEM I
<?)l I
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
4.2.1 Malfunction Description
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.
4.2.2 Malfunction Dynamics
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.
4.2.3 EDS Effectiveness and Crew Safety
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.
4.2.4 Mission Completion Capability
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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2. ENGINES NO. 2 AND NO. 3 AACOAT 145.15 SECONDS
3. VEHICLE FAILS TO MAKE PARKINGORBIT
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F r START TB3 [-START IGMr'S-I 1 90%THRUST140 160 180 260 220
FLIGHT TIME - SECONDS
TIME HISTORY OF DYNAMICVARIABLES FOR
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4-32
D5-15795-7
4.3 ACTUATOR HARDOVER
4.3.1 Malfunction Description
Actuator hardover is any failure that causes an engine actuatorto either extend or retract to its limit.
4.3.2 Malfunction Dynamics
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.
4.3.3 EDS Effectiveness and Crew Safety
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.
4.4.2 Malfunction Dynamics
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.
4.4.3 EDS Effectiveness and Crew Safety
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.
4.4.4 Mission Completion Capability
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
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4.5 SATURATED ERROR SIGNAL
4.5.1 Malfunction Description
Any failure that causes an erroneous 15.3 degree attitude error
signal is a saturated attitude error•
4.5.2 Malfunction Dynamics
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.
4.5.3 EDS Effectiveness and Crew Safety
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.
4.5.4 Mission Completion Capability
Saturated error signal in any stage of flight results in missionloss.
4-46
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4.6 SATURATED RATE SIGNAL
4.6.1 Malfunction Description
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.
4.6.2 Malfunction Dynamics
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.
4.6.3 EDS Effectiveness and Crew Safety
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).
4.6.4 Mission Completion Capability
Saturated rate signal in any stage of flight results inmission loss.
4-51
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D5-15795-7
4.7 LOSS OF INERTIAL ATTITUDE
4.7.1 Malfunction Description
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.
4.7.2 Malfunction Dynamics
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.
4.7.3 EDS Effectiveness and Crew Safety
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.
4.7.4 Mission Completion Capability
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.
4-56
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4.8 LOSS OF ATTITUDE COMMAND
4.8.1 Malfunction Description
Loss of attitude command includes all failures that result in
the attitude command remaining constant or becoming zero atthe time of failure.
4.8.2 Malfunction Dynamics
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.
4.8.3 EDS Effectiveness and Crew Safety
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.
4.8.4 Mission Completion Capability
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
4.9.1 Malfunction Description
Any failure that causes a false attitude error signal of zerodegrees to the flight control computer is referred to as lossof attitude error signal.
4.9.2 Malfunction Dynamics
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.
4.9.3 EDS Effectiveness and Crew Safety
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.
4.9.4 Mission Completion Capability
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.
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:'.i0 LOSS OF ATTITUDE RATE SIGNAL
4.10.1 Malfunction Description
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.
4.10.2 Malfunction Dynamics
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.
4.10.3 EDS Effectiveness and Crew Safety
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.
4.10.4 Mission Completion Capability
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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4-71
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4.11 ACCELEROMETER MALFUNCTIONS
4.11.1 Malfunction Description
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_
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= 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.
4.11.5 EDS Effectiveness and Crew Safety
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.
4.11.6 Mission Completion Capability
_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.
4.12.2 Malfunction Dynamics
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.
4.12.3 EDS Effectiveness and Crew Safety
Aborts are not required for PU malfunctions, and no falseaborts are initiated.
4.12.4 Mission Completion Capability
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
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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
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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.
4.13.2 Malfunction Dynamics
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.
4.13.3 EDS Effectiveness and Crew Safety
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.
4.13.4 Mission Completion Capability
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.
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4.14 LOSS OF BOTH APS MODULES
4.14.1 Malfunction Description
"Loss" is defined as in Paragraph 4.13.1, except both APSmodules are affected.
4.14.2 Malfunction Dynamics
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).
4.14.3 EDS Effectiveness and Crew Safety
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.
4.14.4 Mission Completion Capability
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
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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)
4.15.2 Malfunction Dynamics
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
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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.
4.15.3 EDS Effectiveness and Crew Safety
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.
4.15.4 Mission Completion Capability
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.
4.16.3 Mission Completion Capability
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
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5-10
D5-15795-7
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.
5-11
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DATA NOT AVAILABLE BEYOND THIS POINT.
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(NO S-IVB SECOND BURN)
FIGURE 5-17 0-DEGREE ACQUISITION AND LOSS HISTORY FOR
FOURTH ORBIT FOR Z-ACCELEROMETER MALFUNCTIONS
FOR 78.051-DEGREE LAUNCH AZIMUTH
5-27
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FIGURE 5-22 S-IVB STAGE EARLY SECOND CUTOFF SURVEILLANCE ABO_
0-DEGREE ELEVATION FOR 78.051-DEGREE LAUNCH AZIMUTH
5-32
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