technical report task msc/stl
TRANSCRIPT
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TECHNICAL REPORT
TASK MSC/STL A-20
APOLLO MISSION SA 501
PRELIMINARY MISSION PROFILE (U)
N A S 9 - 2 9 3 8 2 2 F E B R U A R Y 1965
(NASA-CR-103858) APOLLO MISSION S A 501 N79-76 1 4 6
PRELIMINARY 8ISSION PROFILE (TRW Space Technology Labs.) 86 p Unclas
00/12 1 1 1 1 1
TRW SPACE TECHNOLOGY LABORATORIES THOMPSON R A M 0 WOOLORIDGE INC
33 00-600 1 - RCOOO Total Pages:
TECHNICAL REPORT
TASK MSC/STL A-20
APOLLO MISSION SA 501
PRELIMINARY MISSION PROFILE (U)
2 2 F E B R U A R Y 1965
Prepared for
NATIONAL AERONAUTICS A N D SPACE ADMINISTRATION MANNED SPACECRAFT CENTER
Contract No. NAS 9-2938 Phase II (Apollo)
Prepared by &d%.Q*L JW. Dreyfus 0' Task Manager
/
Approved by
(?'j C. V. S tab levd 0 Manager Manned Spaceflight Department
fi / Approved by &?$(pJLTjLpL 3
E. M. Boughton Director Mission Trajectory Control Program
u
TRWSPACE TECHNOLOGY LABORATORIES THOMPSON R A M 0 WOOLORIDGE INC.
R I N T E R V A L S ; R 12 Y E A R S
This document contains information offecting the noti Title 18. U. S. C., Section 793 and 794, the tranrrniss
'ted States within the meanin nner to an unauthorired person is prohibited by low.
3300-600i -RC00G P a g e ii
FOREWORD
This report , which defines the Prel iminary Mission
Profi le fo r Apollo Mission SA-501, is submitted by TRW
Space Technology Laboratories (STL) to the NASA Manned
Spacecraft Center in partial response to Task MSC / S T L
A - 20 (Establishment of Reference Trajectory for Apollo
Mission SA-501) of the Apollo Mission Trajectory Centrol
P rogram (Contract No. NAS9-2938, Phase 11). This repor t
is presented in two volumes. Volume I summarizes the
mission objectives, the system constraints, and the input
data for mission simulation and descr ibes the mission pro-
file. It presents pertinent data in both tabular and graph
forms. Volume I1 contains the t ra jectory listing of the mis-
sion profile, along with the necessary print key.
D
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C O N T E N T S
1 . INTRODUCTION AND SUMMARY . . . . . . . . . . . . . . . . . . . 1.1 Purpose . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.2 Scope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 .3 Profile Summary . . . . . . . . . . . . . . . . . . . . . . . .
2 . SPACECRAFT MISSION REQUIREMENTS . . . . . . . . . . . 2.1 Spacecraft Tes t Objectives . . . . . . . . . . . . . . . . .
2.1.1 First Order . . . . . . . . . . . . . . . . . . . . . . 2.1.2 Second Order ..................... 2.1.3 Third Order ......................
2.2 Mission Constraints . . . . . . . . . . . . . . . . . . . . . . 2 . 2.1 Launch Vehicle System Constraints . . . . . . . 2 . 2 . 2 Spacecraft Systems Constraints . . . . . . . . . 2 . 2 . 3 Trajectory Profile Constraints . . . . . . . . . .
3 . SUMMARY OF INPUT DATA .................... 3.1 Launch Vehicle . . . . . . . . . . . . . . . . . . . . . . . . . 3.2 Spacecraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.3 Ground Stations . . . . . . . . . . . . . . . . . . . . . . . . . 3.4 Miscellaneous Data . . . . . . . . . . . . . . . . . . . . . .
3.4.1 Ear th Constants .................... 3.4. 2 Miscellaneous Constants and Conversion
Fac tors .......................... 4 . MISSION ANALYSIS AND DESCRIPTION . . . . . . . . . . . .
4.1 Saturn V Ascent to Parking Orbit . . . . . . . . . . . . . 4.2 Earth Parking Orbit . . . . . . . . . . . . . . . . . . . . . . 4.3 S-IVB Second Burn . . . . . . . . . . . . . . . . . . . . . . 4.4 Earth Intersecting Coast . . . . . . . . . . . . . . . . . . . 4.5 S P S B u r n . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.6 Pre-ent ry Sequence . . . . . . . . . . . . . . . . . . . . . .
4.6.1 Pre-ent ry Sequence (SPS Burn) . . . . . . . . . 4.6.2 Pre-ent ry Sequence (No SPS Burn) . . . . . . .
4.7 Atmospheric Entry . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 . 7.1
4.7.2
Atmospheric Entry (SPS Burn)
Atmospheric Entry (No SPS Burn) . . . . . . . . 4.8 Vacuum Impact Points ....................
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4 C O N T E N T S (Continued) P a g e
. . . . . . . . . . . . . . . . . . 5. NOMINAL T R A J E C T O R Y DATA 27 6 . TRACKING AND COMMUNICATIONS DATA . . . . . . . . . . 73
R E F E R E N C E S . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75
3300-s001- ,9c000 I P a g e v
ILLUSTRATIONS
b
1-1
3- 1
3-2
3-3
4- 1
4-2
5-1 5- 2
5-3
5-4
5-5
5-6
5-7
5-a
5-9
5-10
5-1 1
5-12
5-13
5-14
Page
Mission Summary . . . . . . . . . . . . . . . . 3
. . . . . . . . . . Saturn V Reference Dimensions 11
Saturn V Zero Angle of Attack Drag Coefficient . . 12 . . . . . Command Module T r i m Drag Coefficient. 13
Effect of S-IVB Second Burn Steering on the Earth Intersecting Coast (Sheets 1 and 2) . . . . . . . . 21, 22
Effect of SM Steering on Entry . . , , , , , , , , 24
Ear th Ground Track . . . . . . . . . . . . . . . Altitude - Longitude History. . . . . . . . . . . . Saturn V Ascent to Orbit/Altitude, Latitude, and Longitude . . . . . . . . . . . . . . . . . . Saturn V Ascent to Orbit/Inertial Velocity, Flight Path Angle, and Azimuth . . . . . . . . . . . . Saturn V Ascent to Orbit/Relative Velocity, Flight Path Angle, and Azimuth . . . . . . . . . . . . Saturn V Ascent to Orbit/Sensed and Total Acceleration . . . . . . . . . . . . . . . . . . Saturn V Ascent to Orbit/Pitch Rate, Pitch Attitude, and Pitch Angle of Attack
Saturn V Ascent to Orbit/Altitude, Mach Number, and Dynamic P res su re . . . . . . . . . . . . . . Ear th Parking Orbit/Altitude, Latitude, and Longitude . . . . . . . . . . . . . . . . . . . Ear th Par-king Orbit /Inertial Velocity, Flight Path Angle, and Azimuth . . . . . . . . . . . . . . Second S-IVB Burn/Altitude, Latitude, and Longitude . . . . . . . . . . . . . . . . . . . Second S-IVB Burn/Inertial Velocity, Flighf Path Angle, and Azimuth . . . . . . . . . . . . . . Second S-IVB Burn/Sensed and Total Acceleration
Second S-IVB Burn/Pi tch Attitude, Pitch Rate, and
. . . . . . . . . . . .
. . Pitch Angle of Attack . . . . . . . . . . . . . .
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ILLUSTRATIONS (Continued)
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5-16
5-1 7
5-18
5-19
5-20
5-21
5-22
5-23
5- 24
5-25
5-26
5-27
5-28
5-29
5-30
5-31
5 - 3 2
5 -33
5-34
Ear th Int e r s e c ting C oa s t /A It itude , La t i tude, and
Ear th Intersecting Coast/Inertial Velocity, Flight Path Angle, and Azimuth. . . . . . . . . . . . . . 48 SPS Burn/Altitude, Latitude, and Longitude . . . . 50
SPS Burn/Inertial Velocity, Flight Path Angle, and Azimuth . . . . . . . . . . . . . . . . . . . 51
SPS Burn/Sensed and Total Acceleration . . . . . . 52
SPS Burn/Pi tch Attitude, Pitch Rate, and Pitch Angle of At tack . . . . . . . . . . . . . . . . . . 53
Pre-en t ry Sequence/Altitude, Latitude, and Longitude . . . . . . . . . . . . . . . . . . . . 5 5
Pre-ent ry Sequence/Inertial Velocity, Flight Path
Pre-entry Sequence (No SPS Burn)/Altitude, Latitude, and Longitude . . . . . . . . . . . . . . 57
Pre-ent ry Sequence (No SPS Burn) /Inertial Velocity,
Atmospheric Entry/Altitude, Latitude, and Longitude - Atmospheric Entry /Inertial Velocity, Flight Path
Atmospheric Entry/Relative Velocity, Flight Path
Atmospheric Entry/Sensed and Total Acceleration 64
Atmospheric Entry /Altitude, Dynamic P r e s sure ,
Longitude . . . . . . . . . . . . . . . . . . . . 47
Angle, and Azimuth. . . . . . . . . . . . . . . . 56
Flight Path Angle, and Azimuth . . . . . . . . . . 58
Angle, and Azimuth. . . . . . . . . . . . . . . . . 62
Angle, and Azimuth. . . . . . . . . . . . . . . . 63
61
and Mach Number . . . . . . . . . . . . . . . . 65 Atmospheric Entry/Range f rom Entry. . . . . . . . 66
and Longitude. . . . . . . . . . . . . . . . . . . 67
Flight Path Angle, and Azimuth . . . . . . . . . . 68
Flight Path Angle, and Azimuth . . . . . . . . . . 69
Atmospheric Entry (No SPS Burn)/Altitude, Latitude,
Atmospheric Entry (No SPS Burn)/Inertial Velocity,
Atmospheric Entry (No SPS Burn) /Relative Velocity,
Atmospheric Entry (No SPS Burn)/Sensed and Total Acceleration . . . . . . . . . . . . . . . . . . . 70
33CS-60G!-RCC00 Page vii
ILLUSTRATIONS (Continued)
Page
5-35 Atmospheric Entry (No SPS Burn)/Altitude, Dynamic P r e s s u r e , and Mach Number . , . , . . . , . . . . 7 1
5-36 Atmospheric Entry (No SPS Burn)/Range F r o m Entry . 72 6-1 Tracking Visibility and Earth Shadow Data , . . . . . 74
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TABLES
Page
3- 1
3-2 3-3
3-4 3- 5 3-6
5-1 5-2 5-3 5-4 5-5 5-6
5-7 5-8
5-9 5-10
Sequential Weight Statement , . , . . , . . . . . . 8
Stage Propulsion Data . . , . . . . . . , . . . . 9 Time Sequence of Events. . . . . . . . . . . . . . 9 Atmospheric Values a t 35 km. . ., . . . . . . . . . 1 2
Spacecraft Weight Data . . . . . . . . . , . , . . 1 3 Radar Tracking Station Sites and Equipment . . . . 14
Time Sequence of Events. . . . . . . . . . . . . . 30 Saturn V Ascent to Orbit/Discrete Events Summary. . 31
Ear th Parking Orbit/Discrete Events Summary. . . . 38 Second S-IVB Burn/Discrete Events Summary . . . , 41
Ear th Intersecting Coast/Discrete Events Summary. , 46
SPS Burn/Discrete Events Summary . . . . , . . . 49 Pre-ent ry Sequence/Discrete Events Summary . . . . 54 Pre-ent ry Sequence (No SPS Burn)/Discrete Events Summary . . . . . . . . . . . . . . . . . 54 Atmospheric Entry/Discrete Events Summary . . . . 59 Atmospheric Entry (No SPS Burn)/Discrete Events Summary . . . . . . . . . . . . . . . . 60
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1. INTRODUCTION AND SUMMARY
1.1 PURPOSE
The Pre l iminary Mission Profile defined in this report is designed
for the unmanned Apollo Mission SA- 50 1.
and spacecraf t t ra jectory profile that is intended to satisfy the miss ion ' s
p r imary spacecraf t objective (to obtain data on the thermal protection
sys tem under lunar entry conditions; Reference 1) without violating any of
the launch vehicle and spacecraft ground ru les and constraints applicable
to the mission.
agreed upon by MSC and MSFC.
1 . 2 SCOPE
It is a combined launch vehicle
It a l so sat isf ies the single-SPS-burn mode previously
In addition to the spacecraf t mission requirements , this volume
of the r epor t summar izes the input data used in simulation of the pro-
file.
analysis for applicable phases, and presents t ime history data for
pertinent t ra jec tory parameters .
and se t t imes as s e e n f r o m 19 tracking stations and the t imes of space-
c raf t entry into and exit f rom the ea r th shadow.
It descr ibes the major phases of the mission, gives the t ra jec tory
It a l so presents the spacecraf t r i s e
Volume I1 of this repor t will contain the ttrajectory listing of the
mission profile.
The computer simulations of both the launch vehicle and space-
craf t t ra jec tor ies were character ized by simple propulsion sys tem
models and open loop steering (constant attitude and attitude ra te
commands).
during orbital coast ope rations and entry, RCS propellant consumption, or
detailed tracking coverage. These i tems a r e currently under evaluation
and will be reported in later documentation on this mission.
These simulations do not consider spacecraft attitude
1.3 PROFILE SUMMARY
Apollo Mission SA-501, current ly planned for the first quarter of
1967, w i l l be the f irst launch of the Saturn V vehicle with an Apollo
spacecraf t . F o r mission simulation, launch is assumed to occur a t
A
3300-6001-RC000 Page 2
13:OO GMT, January 1, f r o m launch complex 39A of the Merr i t t Island
Launch Area. In addition to giving a daylight launch, this time selection
resul ts in a lmost a full day of sunlight near Hawaii f o r the Command
Module (CM) recovery operation.
Major events of the mission a r e i l lustrated in Figure 1-1. The mission has been divided into seven major phases:
1. Saturn V ascent to orbit
2. Ear th parking orbit
3. S-IVB second burn
4. Ear th intersecting coast
5. Service Module Propulsion System (SPS) burn
6. Pre-en t ry sequence
7. Atmospheric entry
The Saturn V launch phase includes the burn of the S-IC stage, the
burn of the S-I1 stage, and a partial burn of the S-IVB stage. Thrust
termination for the la t ter occurs at a 100 n mi circular parking orbit.
After approximately two revolutions in the ear th parking orbit and while
in the vicinity of Cape Kennedy, the S-IVB is res ta r ted and burns to
nominal fuel depletion.
with a n apogee altitude of 7 , 467 n mi and a n inertial flight path angle at
entry into the ea r th ' s atmosphere (defined as 400, 000 feet) of - 7 . 3 5
degrees.
recovery of the CM in case of an SPS failure.
This burn injects the spacecraft into an orbit
This earth-intersecting orbit will permi t successful entry and
Approximately i - 1 1 ' ' 7 L L----- l luula aLLLA C+- - uyv6uu - n f i c r n m a n d ---- while ..___-_ being tracked by
Carnarvon, the SPS is ignited, accelerating the spacecraft so that an
iner t ia l entry velocity of 36, 333 f t / sec and an inertial entry flight path
angle of -7. 35 degrees a re achieved.
sequence is initiated while the spacecraft is still under coverage of
Carnarvon tracking. a f t e r SPS burnout andatmospheric entry occurs about 7 minutes la ter . The
spacecraft then fl ies a Nominal Undershoot Entry Trajectory over a 2500
n mi range and lands approximately 750 n mi due north of the Hawaiian Islands.
Following SPS cutoff, a pre-entry
This coverage is lost approximately 4- 1/2 minutes
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2. SPACECRAFT MISSION REQUIREMENTS
I
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.
2.1 SPACECRAFT TEST OBJECTIVES
The spacecraft tes t objectives presented here were taken f rom
Reference 1.
2 . 1 . 1 First Order
a) Demonstrate satisfactory spacecraft performance a t lunar re turn conditions during terminal t ransear th , entry, parachute-descent, and post-landing mission pha s e s .
b) Demonstrate the s t ructural integrity of the space vehicle (Saturn V/SLA-LEM-CSM).
2 . 1 . 2 Second Order
a) Evaluate CM heat shield performance at the Nominal Undershoot Entry, initiated at not less than the re la - tive velocity corresponding to 36, 333 f t / s ec inertial velocity and an inclination of 40 degrees following a cold-soak to the t ransear th design temperature condition.
b) Determine Emergency Detection System open-loop pe rf o r manc e.
c ) Verify CM radiation shielding effectiveness.
d ) Determine the s t ructural and dynamic response of the CSM and adapter to the Saturn V launch environment.
2. 1 .3 Third Order
Determine the response of the LEM to the Saturn V launch envir.onm ent.
Demonstrate normal mode separation of the Launch Escape System ( L E S I and t h e Boost Protective Cover (BPC) f rom the CSM.
Demonstrate (upon ground command) LES performance in event of launch vehicle failure prior to normal LES jettison.
Demonstrate maximum continuous SPS burn required (approximately 400 seconds to simulate lunar orbit insertion).
Demonstrate operation of the parachute recovery sub- system and recovery aids following entry at nominal design conditions.
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3300-6001-RC000 P a g e 5
f ) Demonstrate operational radiation monitoring instrumen- tation in a radiation environment.
g ) Demonstrate entry guidance at lunar re turn velocity;
2, 2 MISSION CONSTRAINTS
The following mission constraints for this Prel iminary Mission
Profile have been compiled f rom data supplied by MSC and f rom
Reference 1.
2. 2. 1 Launch Vehicle Svstems Constraints
Launch azimuth of 72.0'.
The launch vehicle profile will be a s close as possible to the profile of the nominal Lunar Orbital Rendezvous (LOR) mission.
Full S-IVB loading and full S-IVB burn a r e required.
A minimum of two orbi ts of the S-IVB /IU(Instrumentation Unit)/,%.
Tracking is required for both the pre-ignition sequence and the second S-IVB burn. Eas te rn Tes t Range (ETR) .
This burn will occur over
After S-IVB cutoff, the vehicle attitude will be maintained relative to the local vertical .
Guidance command angle r a t e limitation of 1 deg/sec in pitch and yaw.
Thirty-degree maximum command attitude in the yaw plane.
2. 2. 2 Spacecraft Systems Constraints
a) Total mission duration {launch to Chi spiashj not to exceed 1 2 hours (programmer l imit) .
b) Spacecraft orientation control a s required to provide s t ructural temperature gradients.
c ) A continuous SPS burn of a t least 400 seconds.
d ) Initiation of S M / C M separation manuever no la ter than 5 minutes before reaching 400, 000-foot altitude.
3300-6001 -RC000 Page 6
I
e)
f )
Nominal t r im lift-to-drag rat io is 0. 34.
Tracking is required f o r all SPS burns.
2. 2. 3 Trajectory Profile Constraints
A minimum of two revolutions in the 100 n mi parking orbit.
A t least 4 hours of cold soak beyond ear ths ' significant albedo (to simulate terminal t ransear th conditions) .* F r e e re turn of the spacecraft on an earth-intersecting trajectory following spacecraft/ launch vehicle separation (to allow satisfactory recovery of the spacecraft in case of a n SPS failure to f i re) .
CM entry into the ear ths ' atmosphere (400,000-foot altitude) with a n inertial velocity of 3 6 , 333 f t / s ec and an inertial flight path angle of -7.35 degrees (measured f rom the local horizontal).
A 2500 n mi range f rom entry to landing (Nominal Under shoot Entry Trajectory).
Earth landing to occur in the Pacific Ocean clear of any major island group.
* to SPS ignition is only slightly greater than 3 - 1 / 2 hours. In the selected mission profile, the time duration from S-IVB burnout
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3 . SUMMARY O F I N P U T DATA
The summary of input data in this section consists of data f rom
References 2 and 3 and data agreed upon a t a number of unpublished tech-
nical coordination meetings of MSC and STL personnel.
all quantitative specifications on launch vehicle, spacecraft , and ground
tracking stations, and is considered adequate for the present evaluation
of the mis s ion.
This data includes
3. 1 LAUNCH VEHICLE
Data on the Saturn V launch vehicle w a s based on mater ia l in
Reference 2, and supplemented by an MSFC t ra jec tory listing (dated
2 October 1964) for the SA-501 launch vehicle.
Weight and propulsion charac te r i s t ics of the Saturn V launch
vehicle are presented in Tables 3- 1 and 3- 2, respectively.
are given i n a manner essentially equivalent to their chronological
Weights
disposition in the t ra jec tory simulation.
weights include propellant r e se rve allowances, if any.
is used in a simple propulsion model that applies a constant propellant
flow ra te and a constant thrust , but with correct ions for a tmospheric
p r e s s u r e effects.
t h ree constant-thrust , constant-flow-rate phases to simulate the
optimum thrus t profile for this stage.
occur rence , a re :
It is understood that all jett ison
Propulsion data
The thrust his tory for the S-I1 stage is divided into
These phases , l isted in o rde r of
1) A shor t duration, nominal thrust , nominal specific impulse phase
2) A high thrust , low specific impulse phase
3 ) A low thrust , high specific impulse phase
The launch vehicle sequence of events used in the t ra jec tory simulation
is presented i n Table 3-3 .
weight and propulsion data.
This sequence w a s derived f rom the above
b
At Saturn V Liftoff
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Table 3 - 1. Sequential Weight Statement
Weight Event Losses Weight s
(lb) (1b)
6, 088,000
S-IC Propellant Consumed 4, 192,421
At Inboard Engine Cutoff ’’ S-IC Propellant Consumed
At S- IC Burnout
S- IC at Burnout
At S- I1 Ignition J J *,< .c
S-IC/S-11 Inter stage
LES Jettison:”’
S-I1 Propellant Consumed
At S- I1 Burnout
S-I1 at Burnout
At S-IVB Ignition
S-IVB Propellant Consumed
At S-IVB Cutoff Into Parking Orbit
Loss in Parking Orbit
At S-IVB Second Ignition
S-IVB Propellant Consumed
At S- IVB Burnout
S- IVB at Burnout
1, 895,579
94, 388
I, 801, 191
394, 145
1,407, 046
9, 869
8, 200
929,998
458,979
105, 041
353,938
84, 134
269, 804
3, 873
265,931
141, 312
124, 619
39, 619
1 Payload at Injection 85, 000
.** ‘8.
S-IC inboard engine is cut off 4 sec prior to outboard engine cutoff.
S-IC/S-II interstage and LES a r e jettisoned at 30 and 35 sec af ter S-IB burnout and jettison, respectively.
.L .I. I,. I,.
J d " " 2 Xnn-Anni " " " Z - n r n n n L \ V " " "
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Table 3 -2 . Stage Propulsion Data
Vacuum Sea Level Stage Thrust Thrust
(1b) (1b) s- IC
All Engines Firing 0,745,393 7,610, 000
Only Four Outboard Engine s Firing 6,996,314 6, 088, 000
s- I1 - 0 to 10. 0 sec 1, 035, 000 N/A
10. 0 to 238. 319 sec 1,135, 000 N/A
238. 319 to 367. 072 sec 960,000 N/A
S- IVB 207, 000 N/A
Table 3-3. Time Sequence of Events
Event
Liftoff
End Vertical Rise, Start Pitchover
Inboard Engine Cutoff
Outboard Engine Cutoff and S- IC Separation
S-I1 Ignition, Start High Pitch Rate Steering
Start Low Pitch Rate Steering
Jett ison S-IC/S-I1 Inter stage Adapter
Jett ison Launch Escape System
9-11 Burnout and Jettison, S-IVB Ignition
S-IVB Cutoff Into Orbit
S- IVB Re s t a r t
S-IVB Burnout
Propellant Flow Rate ( lb / sec)
29,496. 14
23,596.91
2,441. 040
2, 694. 234
2, 255.799
488. 206
Time from Liftoff
0 (set)
12,000
142. 135
146. 135
149.935
161.700
176. 135
181. 135
517. 007
689. 340
T + O
T -f- 289.451
C
3 300- 600 1 - R G O U U Page 10
The Saturn V launch vehicle is i l lustrated in Figure 3-1, and the
An ze ro angle-of-attack drag force data is presented in Figure 3-2.
aerodynamic reference a r e a of 855.3 square feet was used.
The static atmosphere model used in the t ra jec tory simulation has
Below an altitude of 35 km, the Pa t r ick Atmosphere (Refer- two par ts .
ence 4) is used, while between 35 km and 400,000 feet, the U.S. Standard
Atmosphere, 1962 (Reference 5) is used. Both atmospheres a r e provided
to the computer a s tables of p re s su re and tempera ture versus geometric
altitude. In changing f rom the Pa t r ick to the U.S. Standard Atmosphere
at a n altitude of 114,830 feet , no attempt has been made to remove the
discontinuity between the two models.
a tmospheric values at that altitude.
Table 3 - 4 gives comparative
3 . 2 SPACECRAFT
Weight charac te r i s t ics for the Apollo spacecraf t were obtained from
Reference 6. en t ry aerodynamic data were obtained in unpublished technical coordination
meetings with MSC personnel.
spacecraf t is presented in Table 3-5.
SPS propellants a re not consumed, which should provide more than adequate
allowance for flight performance r e se rves .
SPS thrus t and propellant flowrate charac te r i s t ics and CM
The sequential weight statement for the
Approximately 1400 pounds of usable
The SPS w a s character ized by a vacuum thrus t of 21,900 pounds
and a vacuum specific impulse of 313. 0 seconds.
propellant flow rate of 69. 968 lb/sec.
This resul ts in a
The aerodynamic t r i m drag coefficient data used for the CM ent ry
t r a j ec to ry is presented in Figure 3 - 3 and is based on an aerodynamic
referezce area nf 139. 36 sqi~are feet.
constant 0. 34 t imes the d r a g force.
the U. S. Standard Atmosphere, 1962, and a constant CM weight of
11,000 pounds.
The l i f t force is based on a
The en t ry t ra jec tory is based upon
3. 3 GROUND STATIONS
The prec ise complement of stations, their ultimate locations,
the i r operating charac te r i s t ics , and such related data to be used for
Apollo Mission SA-501 a r e currently not known.
compiled f rom tracking site and equipment information supplied by
MSC, is assumed to be applicable at this time.
The data in Table 3 - 6 ,
3300-600 I-RC000 Page 11
STAT I ON S
MA1 N STAGE PROPELLANT CAPAC I TY 930,000 LB
MSFC DWG 0299 REV.C
Figure 3- 1. Saturn V Reference Dimensions
3300-600 1 -RC000 I Page 12
7,938 1 + 1,000M + ‘D,~*REF
D =
.o
MACH NUMBER
Figure 3-2. Saturn V Zero Angle of Attack Drag Coefficient
Table 3-4. Atmospheric Values at 35 k m
Patr ick Atmosphere U. S. Standard
0. 085922154 0. 08341 8396 2 P r e s s u r e (lb/in )
Temperature (OR) 439.42235 425.72477
Speed of sound (f t /sec) 1027. 6009 1011.4798
Density (slug/ft ) 0. 16403886 x 0. 1 6 4 3 7 5 8 ~ 1 0 - ~ 3
c
3 3 00 - 600 1 -RCO 00 P a g e 13
Table 3-5. Spacecraft Weight Data
Weight Losses
(1b) At Liftoff
Launch Escape Sys tern 8 , 2 0 0 At Injection Into Orbit
Event Weights
(lb) 93, 200
8 5 , 0 0 0 LEM Jet t ison Saturn Launch Adapter
At SPS Ignition
21,490 3 ,800
59,710
SPS Propel lant Consumed 37, 083* At SPS Cutoff
SPS Propellant Remaining 1,427 SPS Burnout 10,200
22,627
Command Module 11,000
Propel lant consumption is based on a 530-second SPS burn time. 'P
1.4
1.3
1.2 UJ V - LL LL
5 c3 25 n V 2 1.0 Q z i n 0 5 0.4 4
1.1
-
0.8
0.7
MACH NUMBER Figure 3-3. Command Module Tr im Drag Coefficient
3300-600 1 - R C 0 0 0 Page 14
- 0 m VI a
- 0 Wl
v3
v1 a 5
a d rd 4 (0 H
k 0 a rd > rd VI d
4
4
d a 5 k a, a
rd
rd k u rn
.rl 4
4 n
rd k k a, P d rd u
0 V .d
B - rn rd
b d E
3
3300- 6001 - ,9co00 Page 15
The station coordinates given a r e based on a Fisher ellipsoid.
This model is described by:
a = semimajor axis = 6378. 166 km
b = semiminor axis = 6356.784 k m
f = flattening = 1/298. 3
The altitude is referenced to the ellipsoid and includes geoidal separation.
3 . 4 MISCELLANEOUS DATA
The following ear th constants and conversion factors (Reference 7 )
have been used in the generation of the mission profile and are consistent
with those presented in Reference 8.
3. 4. 1 Ear th Constants
Rotational rate 4.37526902 x 1 0'3 rad /min
0.417807416 x deg/sec
0. 729211504 x r ad / sec
Equatorial radius
Average radius
Gravitational parameter he)
2. 092573819 x 107 f t
2. 0909841 x 107 f t
5. 53039344 x 10'3 e r3 /min2
11. 46782384 x 103 er3/day2
3.986032 x 105 km3/sec2
1. 407653916 x 1016 ft3/sed2
Coefficients of potential harmonics
J term (second harmonic)
H term (third harmonic)
D term (fourth harmonic)
1. 62345 x 1 0'3 nd
-0. 575 x 10-5 nd
0.7875 x 10-5 nd
,
Ear th flattening (f) 1 /298. 3 nd
3 3 0 0 - 600 1 - RCOOO Page 16
3. 4. 2 Miscellaneous Constants and Conversion Fac tors
Velocity of light in a vacuum
Astronomical unit of length
Kilometers p e r foot
9. 835711 94 x 1 O8 f t / s ec
4. 9081 0367 x 1 0l1 f t
0. 3048 x 1 0'3 km/f t
Kilometers p e r nautical mile 1.852 km/n mi
Fee t p e r nautical mile 6076. 115486 f t / n mi
Weight - to-mass rat io 3 2. 17404856 lb/ slug
Mass -to-weight ra t io
Fee t p e r ear th equatorial radius
0. 031 080950 s lug/ lb
2. 092573819 x 1 O7 f t /er
Nautical mile pe r ear th equatorial radius
3443. 93358 n mi/er
3500-6001 -RC000 Page 17
4. MISSION ANALYSIS AND DESCRIPTION
The miss ion profile f o r Apollo Mission SA-501 has been designed
to meet the Tes t Objectives of Section 2. 1 without violating the Mission
Constraints of Section 2.2. To sat isfy these objectives and constraints
and determine values of the f r e e variables, a certain amount of t ra jectory
analysis was required. The results of this analysis, along with a descrip-
tion of the resulting mission profile, a r e given in this section.
In addition to the nominal mission profile, an alternate profile resul t -
ing f rom an SPS fai lure to burn is also described. These two profiles have
been characterized by "SPS Burn" and "No SPS Burn" titles and a r e identi-
ca l up to the SPS ignition.
In the selected profile, the time duration f rom S-IVB burnout to
SPS ignition is slightly greater than 3-1/2 hours and therefore does not
m e e t the 4-hour cold soak constraint (Section 2 .2 . 3b).
4 .1 SATURN V ASCENT TO PARKING ORBIT
Launch of Apollo Mission SA-501 will occur from Pad "A" of Launch
Complex 39 during the f i r s t quarter of 1967.
a r e 28O38' 50. 927" North latitude and 80°38! 08.0711' West longitude.
the t ra jectory simulation, launch was assumed to occur at 13:OO hours
GMT (08:OO hours EST) on 1 January 1967.
The geodetic coordinates
F o r
The launch profile is initiated with a 12-second vertical r i s e
followed by a 0. 3568-degree kick (an instantaneous rotation of the missile
attitude and velocity vector) into a gravity turn t ra jectory with a 72-
degree azimuth heading.
board engine cutoff, the center engine is cut off.
coas t f r o m S-IC cutoff and separation, the S-I1 is ignited and initiates a
pitch-up a t a 1 deg/sec rate.
approximately 11. 5 seconds after ignition, and a low pitch rate s teer ing
of 0. 1043 deg/sec down is initiated.
and separation, the S-IC/S-I1 interstage adapter is jettisoned, and 5 seconds l a t e r , the LES is jettisoned.
Approximately 4 seconds pr ior to S-IC out-
Following a 3. 8-second
This high pitch rate steering is terminated
Thirty seconds after S-IC cutoff
3300-6001 -RC000 Page 18
i
The S-I1 burnout and separation and the S-IVB ignition all occur
simultaneously in the simulation pe r Reference 2.
s teer ing of 0. 1043 deg/sec is maintained until S-IV B cutoff at c i r cu la r
orbi t velocity.
The low pitch rate
The values of the kick angle, the duration of the high pitch rate
s teer ing, and the magnitude of the low pitch rate steering were
determined by i terat ion techniques so that the S-IVB cutoff would occur
at 100 n mi altitude with a zero-degree flight path angle and maximum
injected weight (or concurrently, minimum S-IVB propellant
consumption).
It should be noted that the injected weight of 269, 804 pounds calcu-
la ted in the above simulation is 366 pounds l e s s than that presented in
Reference 2. This amounts to 0. 136 percent of the injected weight, o r
l e s s than 1 second of S-IVB burning, and is well within the e r r o r to be
expected due to non-nominal performance.
for establishing the prel iminary miss ion profile and the spacecraf t launch
environment.
These resul ts a r e sat isfactory
4 .2 EARTH PARKING ORBIT
The S-IVB cutoff, which occurs approximately 11. 5 minutes af ter
liftoff, injects the spacecraf t into a 100 n mi c i rcu lar parking orbi t with
a n inclination of 32. 588 degrees and an orbital period of 88.1 minutes.
On the second orbit , approximately 169 minutes af ter S-IVB cutoff, the
Point Arguello Tracking Site acquires the spacecraf t a t a 5.0-degree
elevation angle f r o m the horizon, in preparation f o r the S-IVB second
burn ignition. This initial acquisition by Point Arguello marks the begin-
ning - of a period of continual tracking coverage over the continental United
States.
4.3 S-IVB SECOND BURN
Eleven minutes a f t e r Point Arguello tracking acquisition, the space-
c raf t has passed over the Eastern coast and is approximately 240 n m i
north of Cape Kennedy. At this point, the S-IVB is ignited and starts its
second burn.
t h ree reasons:
This particular selection of ignition time was chosen for
c
3300-6001 -RC000 P a g e 1 9
It is desirable to have a good period of tracking coverage pr ior to S-IVB ignition. The above selection allows 11 minutes or tracking coverage.
It is desirable to have a good period of tracking coverage following S-IVB burnout. This selection allows a 12- minute period of Antigua tracking coverage f rom burnout to the time that the spacecraf t goes below a 5. 0-degree tracking elevation angle.
F o r the ear th intersecting coast profile selected, this timing places the CM splash point due north of the Hawaiian Islands. point, for it places the impact position of the spent S-IVB stage and the SM near the middle of the Pacific Ocean and is a convenient position fo r operating the recovery force.
This appears to be a desirable landing
The S-IVB burns for a full 290 seconds, simulating a t ranslunar
injection. However, the pitch steering w a s determined so that the
result ing orbit would remain near ear th (less than 12,000 n mi altitude)
and in te rsec t the ea r th ' s atmosphere.
CM recovery i f the SPS fails to fire.
The la t te r allows fo r a successful
At this date, there a r e no firm c r i t e r i a on which to base the
selection of an entry flight path angle for the no-SPS-burn condition.
T h e r e appear to be two alternatives, however.
per form a high heat ra te tes t of the CM i f the SPS fails to burn.
value of the entry flight path angle is required in this case since the
entry velocity is considerably less than the desired value of 3 6 , 3 3 3 f t /sec.
However, this alternative has an operational disadvantage in that the
locations of atmospheric en t ry and CM splash f rom the no-SPS-burn
case are considerably separated f rom those with an SPS burn, and two recoxrery team-s wnuld probably be needed to cover both landing locations.
The first would be to
A large
3300-6001 - R C O O O P a g e 20
The other alternative would be to select a no-SPS-burn entry
condition in which the distance between the two landing sites would be
minimized. This requirement would favor a shallow entry flight path
angle in o rde r to s t re tch the entry range to a maximum.
alternative was developed in this mission profile, and a flight path angle
of -7. 35 degrees (same as that required wi th an SPS burn) was selected.
This results in a 3 6 3 n mi separation of the two CM splash points.
The la t te r
::< The effect of the S-IVB pitch ra te on the initial pitch attitude
required to achieve entry flight path angles of -7. 35 , -10.0, -12 . 5 , and
-15.0 degrees is i l lustrated in Figure 4-1, Sheet 1.
shows the resulting apogee altitudes and entry latitudes. F igure 4-1, Sheet 2 ,
shows the effect of S-IVBpitch rate on entry longitude, the t ime duration
from liftoff to entry , and the entry velocities,
tions concerning pitch steering may be made f rom this data:
This illustration a l so
Several interesting observa-
Large values of S-IVB pitch rate, both positive and negative, and grea te r in magnitude than 0. 3 deg/sec , appear desirable since they place the atmospheric entry into the Pacific Ocean and c l ea r of the Asian continent, and secondly, reduce the total mission duration time.
Conversely, values of S-IVB pitch rate l e s s than 0. 3 deg/ sec in magnitude have ve ry long mission durations, very poor locations of atmospheric entry location, and the highest entry velocities.
Negative pitch ra tes (counterclockwise and up) require negative (down) values of the init ial pitch attitude. is deemed undesirable because i t will cause the altitude to decrease initially and is an und.esirable thrusting attitude.
This
Based on the above observations, an S-IVB pitch rate of 0. 5 deg/
s ec was selected f o r the second hnrn.
attitude of - 122.4 degrees (up) for the entry flight path angle of - 7 . 3 5 degrees.
This r z s d t s in 211 iiiiiiai pitch
... -8-
F o r simulation purposes, the initial attitude is referenced to the iner t ia l velocity vector.
9 0
w. I
0 *
0 rn
n
W 0 n
0 - W W
n 3 I- + 4 2 s + W Z
0
0 I - 0 w.
Z v
0
0 (u
I
0 * 7
9'; I
O N O N
* 9. 0
I 0
t 0
I
0 cu. 0 I
cu. 0
t 0
(33S/E>3a) 31VY H311d fla-S
I 9. 0 I
3300-6001 -RC000 Page 23
Unfortunately, this selection does not meet the 4-hours cold-soak constraint of Section 2. 2 . 3 . b.
is only 3 hours and 34 minutes with this profile.
The elapsed time from S-IVB burnout to SPS ignition
4 . 4 EARTH INTERSECTING COAST
At final burnout, 3 hours and 16 minutes af ter liftoff, the S-IVB
has injected the spacecraft into an orbit characterized by a 7,467 n mi
apogee altitude and a -7. 35-degree entry flight path angle. This orbit
has a semimajor axis of 43, 573,000 feet , an orbital eccentricity of
0. 5226, and an orbital inclination of 32.6 degrees.
Although an exact coast duration from S-IVB burnout to S-IVB/CSM
separation was not established in the launch vehicle constraint ( 2 . 2 . I f ) ,
a 30-second coast period was assumed for this profile. Approximately
12 minutes af ter S-IVB burnout, the spacecraft goes below the 5-degree
elevation angle f rom the Antigua tracking station.
coverage over the continental United States before, during, and after the
S-IVB burn should provide adequate ground tracking data for launch
vehicle systems evaluation and spacecraft orbit determination.
The ground tracking
Almost an hour and a half la ter , Carnarvon tracking observes the
CSM a t an elevation angle of 5 . 0 degrees and a range of 9500 n mi.
Twenty-three minutes la te r ( 5 hours and 15 minutes af ter liftoff), the
spacecraf t reaches apogee.
but p r io r to SPS ignition, the orbit state vector (position and velocity) and
the target vector will be updated in the Apollo Guidance Computer (AGC)
by Carnarvon tracking via the Up-Data -Link.
During the spacecraft descent from apogee,
4 .5 SPSBURN
The nominal mission plan calls for the SPS to accelerate the
spacecraf t in such a way that the desired values of velocity and flight
path angle at atmospheric entry wi l l be achieved.
the effects of SPS pitch rate and SPS ignition time (measured f rom
apogee) on the SPS initial pitch attitude that is required at ignition in
o r d e r to achieve the -7. 35-degree entry flight path angle.
sponding values of the entry velocity and time duration f rom SPS burnout
t o atmospheric en t ry are a l s o shown in this figure. This data indicates
that the entry velocity achieved f rom the 530-second SPS burn continues
t o increase as the SPS ignition time is delayed.
Figure 4-2 i l lustrates
The co r re -
0 9
0 rn
0 *
0 n
0 hl
0 c-
0
r i i
c? 0
cu. 0
U 9 7
0 - . 0 - cu.7 9 0
I 0
The selection of a 45-minute period from apogee to ignition appears
to be a reasonable one based on the above data. This selection allows
for an 11, 3-minute period from SPS burnout to atmospheric entry.
Resulting values of SPS pitch rate and initial pitch attitude
3 6 , 3 3 3 f t / s ec entry velocity and the -7. 35-degree entry flight path angle
were -0. 292 deg/sec (counterclockwise) and t 1 2 9 . 3 degrees (down),
r e s pe c tive ly.
* to satisfy the
4 . 6 PRE-ENTRY SEQUENCE
4. 6. 1 Pre-Ent ry Sequence (SPS Burn)
Almost 7 hours after liftoff, the SPS is cut off and an 11. 3-minute
At 3. 8 minutes after SPS cutoff, coast to atmospheric entry is started.
the spacecraft drops below a 5. 0-degree Carnarvon tracking elevation
angle,
separation is initiated. The CM then assumes the proper entry attitude.
Atmospheric entry occurs a t a longitude of 155. 64" East and a latitude
of 23. 40" North.
This profile assumes that 5 minutes pr ior to entry, CM/SM
4. 6. 2 Pre-Ent rv Seauence (No SPS Burn)
The pre-entry sequence phase i n case of SPS failure to burn is
initiated a t the nominal SPS ignition t ime, 45 minutes af ter apogee.
Approximately 5. 3 minutes of Carnarvon tracking at elevation angles
greater than 5. 0 degrees remain after this time, and atmospheric entry
occurs 12. 1 minutes after loss of Carnarvon tracking. Entry occurs at
170. 85" East longitude and 28. 95" North latitude with an iner t ia l velocity
of 31,592 f t / s e c and a n iner t ia l flight path angle of -7. 35 degrees.
before, CM/SM separation is assumed t o occur 5 minutes pr ior to entry.
As
4. 7 ATMOSPHERIC ENTRY
4. 7. 1 Atmospheric Entry (SPS Burn)
Seven hours and 18 minutes after liftoff, the CM initiates entry into
the ear th ' s atmosphere.
two values of the ver t ical lift-to-drag ratio.
w a s used f r o m ent ry to pullout (horizontal flight) while a lower value
The entry t ra jectory was simulated by using
The maximum value 0. 3 4
* Referenced to the iner t ia l velocity vector.
3300 - G O O 1 -RC000 P a g e 26
was used from pullout to drogue chute deployment a t 24,000 feet.
s ea rch i teration w a s performed on the la t ter lift-to-drag ratio in order
to sat isfy the 2500 n mi entry range requirement.
drag ratio of 0.2346 simulates the effect of the Apollo CM rolling back
and forth to reduce the ver t ical lift component.
A
The resulting l if t- to-
Horizontal pullout is achieved 76. 3 seconds after entry, and the
CM starts an upward drift at the reduced lift-to-drag ratio.
ascends to an altitude of 360, 400 feet before it starts the final descent
5 -1 / 2 minutes after atmospheric entry.
9 minutes la te r a t an altitude of 24,000 feet .
parachute deployment a t an altitude of 11 , 000 feet , the CM performs a
water landing almost 16 minutes after entry and almost 7-1 / 2 hours
a f t e r liftoff. The splash point latitude and longitude are 32. 46" nor^
and 157. 98" West, respectively. This position places the spacecraft
approximately 750 n mi north of Hawaii.
The CM
Drogue chute deployment occurs
Following the main
4. 7. 2 Atmospheric Entrv (No SPS Burn)
The entry t ra jectory profile for the no-SPS-burn condition has
a sequencing of events identical to that presented above. The pr imary
difference in the two profiles is that the entry velocity for the no-SPS-
burn is 5740 f t / s ec l e s s than desired.
this profile, a range of 1254 n mi f rom entry to landing can be achieved.
The water landing occurs 11. 4 minutes after en t ry and almost 7-1 / 2 hours
a f te r liftoff at 32. 54" North lati tude and 165. 15" W e s t longitude.
sp lash point is approximately 363 n mi f rom that estimated fo r the
nominal, SPS -burn profile.
By using maximum l i f t throughout
This
4. 8 VACUUM IMPACT POINTS
Impact points for the S-IC, S-11, and S-IVB have been calculated,
based upon a vacuum ballistic entry.
SM since burnout inser t s the SM into an orbit with a perigee altitude of
1 0 n mi.
No impact was available for the
S-IC impact occurs at 30. 16" North latitude and 74. 59" West
longitude while S-11 impact occurs at 32. 05" North latitude and 38. 32"
West longitude.
at 31. 48" North latitude and 176. 90" West longitude.
The S-IVB impact following the second burn will occur
5. NOMINAL TRAJECTORY DATA
This section contains trajectory parameter histories describing and
illustrating the nominal mission profile. These data, presented here in
tabular and graph f o r m s , a r e based on the t ra jectory printout data in
Volume I1 of this report . Data a r e presented for both the SPS burn profile
and the no-SPS-burn profile,
Figures 5-1 and 5-2 present the ear th ground t rack and the altitude-
longitude history, respectively, for the entire mission profile. The t ime
sequence of events for the mission is shown in Table 5-1.
F o r each of the mission 's seven major phases, pertinent powered and
and f r e e flight trajectory parameters have been plotted a s a function of
time f rom liftoff. These graphs, along with related tabular data, have
been grouped on the following pages according to mission phase, as follows:
1) Saturn V Ascent to Orbit (Table 5-2, Figures 5-3 through 5-8)
2 ) Ea r th Parking Orbit (Table 5-3, Figures 5-9 and 5-10)
3 ) Second S-IVB Burn (Table 5-4, Figures 5-11 through 5-14)
4) Ear th Intersecting Coast (Table 5-5, Figures 5-15 and 5-16)
5 ) SPS Burn (Table 5-6, Figures 5-17 through 5-20)
6 ) P re -en t ry Sequence (Tables 5-7 and 5-8, Figures 5-21 through 5 - 24)
7) Atmospheric Entry (Tables 5-9 and 5-10, Figures 5-25 through 5-36)
0 9 8 0
0 9
f
o z o
0 0 , T 'p I
3
I 1 I
1 I I
W I
W e p. I-
2 5
n (3
W n
(3 Z 0
2 -
A
I rn I 1 I I
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(,01 x i 4 3aniiiiv
c
u4 S
I
n
W (3
n v
w n
(3 Z 0
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J
P) a 5 c, .d
? s I
Table 5-1. Time Sequence of Events
Sa tu rn V A R C C I I ~ to park in^ Orbi t
Liftoff End V r r t i r a i R i se , S t a r t P i tchovcr S - IC Cc*ntcr Enginr Cutoff S- IC Outboard Enginc Cu to f f , S- IC Srpa ra t ion S-11 lgnition s - l C / s - I I In t e r s t age Adapter Je t t inon Launch E ~ r a p c Sys t em Je t t i eon S-11 Engine Cu to f f , S-I1 Je t t ieon , and S-IVB Ignition S-IVD Engine Cu to f f ln to E a r t h Pa rk ing O r b i t
E a r t h P a r k i n g Orb i t
S t a r t of E a r t h P a r k i n g O r b i t S t a r t of Second Orb i t Acquinit ion of Po in t Arguel lo Track ing End of E a r t h P a r k i n g Orb i t
Second S-IVB B u r n
S t a r t of S-IVD Second B u r n Burnout of S-IVB
E a r t h ln t c rnec t ing C o a s t
S t a r t of Coas t Loa Antigun Tracking S- IV B Je t t i s on Acqure C a r n a r v o n Track ing Apogee of Coas t ( 7 , 4 6 7 n m i Altitude) Update Spacec ra f t S t a t e Vec to r (Pos i t i on and Velocity) End of E a r t h In t e r sec t ing C o a s t
SPS B u r n
S t a r t of SPS B u r n S P S Cutoff
Pr e - e n t r y Sequenc c
S t a r t of C o a s t to Cn t ry Lose C a r n a r v o n Track ing SM Je t t i son and Aesume En t ry Attitude 4 0 0 . 0 0 0 ft Alti tude
At rno e ph e r i c E n t r y
S t a r t of E n t r y T r a j e c t o r y Pul lout to Hor izonta l F l igh t , S t a r t R o l l Maneuvering Drogue Chute Deploymcnt a t 2 4 , 0 0 0 f t Main Pa rachu te Deployment a t 1 1 , 0 0 0 f t E a r t h Landing
Q No SPS B u r n P ro f i l e
P r e - e n t r y Sequence
S t a r t of Coae t to E n t r y LoRe C a r n a r v o n Track ing SM Je t t i son and Assume E n t r y Attitude 400,000 I t
Atmonpher ic En t ry
S t a r t of F:lltry T r a j c r t o r y Pul lout LO Ilorinontal Fl ight . S t a r t Roll Mancuvering 1)rogc Chrltc! J)cploymcnt at 2 4 , 000 It Meiin Chutc 1)eploymt:nt ;it 1 I , 000 f t E;irth 1,;indiiig
Tirnc: F r o m Liftoff (li r : mi t i : e c! c , GM T ) _-
1 3 : O O : O O I3:OO: 12 .00 13:OZ:ZZ . 1 3 13:02:26. 13 13:02: 29:93 13:02:56. I 3 1 3:03:0 1 . 1 3 13:08:37.01 1.3:11:29.34.
13: 11:29. 34
16:OO: 3 7 . 6 9 16: I I : 3 7 . 6 9
- 14:38:00
\ 16: 11 :37 .69 16:16:27. 14
16:16:27. 14 16: 28: 3 0 . 0 0 16:16:57. 14 17:52:02.20 18: 15: 17. 32 19:36:00.00 19:50:37. 3 3
19:50:37. 3 3 19:59:27. 33
19:59:27. 33 20:03: 15..28 20:05:40.00 20: 1 0 : 4 9 . 9 3
20: 10:49.93 20: 12 :06 .29 20:25:32.65 20:26:06.27 20:26:40. ea
19:50:37. 33 20 :05:56 .73 20:12:00.00 20: 17 :57 .63
20 :17:57 .63 20: 1 9 : 2 2 . 23 20:28: I I . 29 20 :28:46 .65 Z O : Z 9 : 2 2 . 94
*The evc:nt 1ilrlc-H .ire the! Hiiiiic through thc Ear th In tc recc t ing Coant phanc.
33oo-buu l - l~c ;ooo Page 31
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3300-600 1 -RCOOO Page 32
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4 .0 30.5 x lo5
3 . 0 30.0 x lo5
2 .0 29.5 x lo5
1.0 29.0 x lo5
0.0 28.5
TIME FROM LIFTOFF (SEC)
Figure 5-3. Saturn V Ascent to Orbit/Altitude, Latitude, and Longitude
3300-6001 -RC000 Page 33
5 . 0 45 .0 x lo4
4 .5 40.0 x lo4
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Figure 5-4. Saturn V Ascent to Orbit/Inertial Velocity, Flight Path Angle, and Azimuth
5.0 90.0 x lo4
4.5 80.0 x lo4
4.0 70.0 x lo4
3.5 60.0 x lo4
1.5 20.0 x lo4
1.0 10.0 x lo4
0.5 0.0 x lo4
0.0 -10.0
3300-6030 i-RC000 Page 34
TIME FROM LIFTOFF (SEC)
Figure 5-5. Saturn V Ascent to Orbit1Relative Velocity, Flight Pa th Angle, and Azimuth
3300-6001-RC000 Page 35
TIME FROM LIFTOFF (SEC)
Figure 5-6. Saturn V Ascent to Orbit/Sensed and Total Acceleration
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Figure 5-7 . Saturn V Ascent to Orbi t /Pi tch Rate, Pi tch Attitude, and Pi tch Angle of Attack
3300-6001-RCOOC Page 37
10.0 10.0 1ooo.c x lo5
9.0 9.0 900.0 x lo5
8.0 8.0 800.0 x lo5
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Figure 5-8. Saturn V Ascent to Orbit/Altitude, Mach Number, and Dynamic P r e s s u r e
3300-6001 -1ZC.000 Page 3 8
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3300-600 1-RC000 Page 40
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3300-60bl-RC000 Page 42
Figure 5- 11 . Second S-IVB Burn/Altitude, Latitude, and Longitude
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3300-6001-RCOOO P a g e 43
TIME FROM LIFTOFF (SEC) x lo4
F i g u r e 5- 12 . Second S-IVB B u r n / I n e r t i a l Veloc i ty , F l i g h t Path Angle, a n d A z i m u t h
t v)
3300-6001-RCOOO Page 44
TIME FROM LIFTOFF (SEC) x lo4
Second S- IVB Burn/Sensed and Total Acceleration Figure 5- 13.
3300-600 1 -RCOOU Page 45
1 .of-
0.8
0.6
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120
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TIME FROM LIFTOFF (SEC) x lo4
Figure 5- 14. Second S-IVB Burn/Pi tch Attitude, Pitch Rate, and Pitch Angle of Attack
3 3 0 0 - 60 0 I - !?. c; n 0 0 P a g e 46
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3300-600 1 -RC000 Page 47
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Figure 5- 15. Ea r th Intersecting Coast/Altitude, Latitude, and Longitude
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3300-6001-RCOOO Page 48
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TIME FROM LIF~OFF (SEC) x lo4
Figure 5- 16. E a r t h Intersecting Coast/Inertial Velocity, Flight Pa th Angle, and Azimuth
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3300-600 1-RC000 Page 50
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3300-600 1 -RC000 Page 55
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Figure 5 - 2 1. Pre -en t ry Sequence/Altitude, Latitude, and Longitude
3300-600 1 -RC000 Page 56
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Figure 5 - 2 2 . Pre-en t ry Sequence/Inertial Velocity, Flight Path Angle, and Azimuth
33OO-6OOl-RCOOO Page 57
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3300-600 1 -RC000 Page 58
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3300-600 1-RC000 Page 66
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3300-6001-RC000 Page 67
F igure 5- 31. Atmospheric Entry (No SPS Burn)/Altitude, Latitude, and Longitude
3300-6001-RCOOO Page 68
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Figure 5-32. Atmospheric Entry ( N o SPS Burn) /Inertial Velocity, Flight Pa th Angle, and Azimuth
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3300-600 1 -RCOOO Page 69
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Figure 5- 3 3 . Atmospheric Entry (No SPS Burn)/Relative Velocity, Flight Path Angle, and Azimuth
3300-6001-RC000 Page 70
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3300-600 1 -RCOOO Page 71
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3300-600 1 -RC000 Page 72
\
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6. TRACKING AND COMMUNICATIONS DATA
Spacecraft visibility periods for the tracking stations presented
in Table 3 - 6 a r e illustrated in Figure 6-1.
defined as a tracking elevation angle greater than 5. 0 degrees.
F igure 6-1 a l s o i l lust rates the t ime periods when the spacecraft is in the
ea r ths ' shadow.
Spacecraft visibility is
It is expected that at least one injection tracking ship and one entry
t racking ship wil l be available for this mission. Visibility data is not
presented for these stations because the i r placement for this mission
has not been established.
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3300-6001 -RC000 Page 75
REFERENCES
1.
2.
3.
4.
5.
6.
7.
8.
9.
"Objectives, Requirements, and Flight Profile for Missions 501 and 502, MSC/ASPO/Test P rogram Planning Branch, 28 September 1964.
"Minutes of Sixth Meeting of the Guidance and Performance Sub- Panel (U)ll (Enclosure 8 titled "Prel iminary Saturn V SA-501 Trajectory and Design Trade Studies"), MSFC/MSC , 3 0 October 1964. (C)
"Lunar Orbit Rendezvous Reference Trajectory Data Package - Issue 4 (U), I t TRW Space Technology Laboratories Report No. 8408-6056-RC-000, 8 July 1964. (C)
0. E. Smith, "A Reference Atmosphere for Patr ick AFB, Florida, 'I NASA Technical Note D 595, March 1961.
"U. S. Standard Atmosphere, 1962, U. S. Government Printing Office, Washington, D. C. , 1962.
"Minutes of Fourth Meeting of the Guidance and Performance Sub- Panel (U), MSFC/MSC, 10 September 1964. (C)
J. 0. Cappellari, Jr. , "Standard Astrodynamic Constants and Conversion Factors for Project Apollo, 8 January 1964.
Bellcomm Inc. ,
"Apollo Operational Nominal Trajectory Ground Rules, I t MSC Internal Note No. 64-OM-4, 1 4 March 1964.
"Design Reference Mis sion-Apollo Mission Planning Task Force, GAEC Report No. LED-540-12, 3 0 October 1964.
DISTRIBUTION
J. P. Mayer, IITASA/L4SC/IG"JP (225 t 1 reproducible)
NASA/IGC Technical Information Division (4)
2. C. Liounis, PASA/MSC Control Systems Procurement (1)
TRWISTL (Apollo Distribution + 2 reproducibles )