:’ -.

This document is prepared jointly by the Marshall Space Flight Center Laboratories R-ASK-P, R-ASTE-S, and R-P&"E-VN. The document presents a brief and concise description of the AS-501 Apollo Saturn . Space Vehicle. Where necessary, for clarification, addirional related information has been included.

It is not the intent of this document to completely define the Space Vehicle or its systems and subsystws in detail. The infomacian presented herein, by rext and sketches, describes launch preparation activities, launch facilities, and the space vehicle. This information permits the reader to follow the space vehicle sequence Of events begin- ning a few hours prior fo liftoff to its journey into space.

1. Mission Purpose:

The purpose af the AS-501 mission is to develop the Saturn V launch vehicle for manned flights and to verify the adequacy of the Apollo Command Module heat shield at lunar reentry velocities.

The AS-501 mission is an unmanned, elliptical earth orbital flight.

2. Mission Objectives:

a. Demonstrate structure and thermal integrity, and cmpati- biliry of the launch vehicle and spacecraft, and determine structural loads and dynamic characteristics during powered and coasting flight.

c. Verify launch supporr equipment compatibility, and mission support capability for launch and mission operations to high post-injec- tion altitudes and Command Module recovery.

d. Demonstrate the S-XC and S-II stage propulsion systems and * determine in-flight system performance parameter.

e. Dwonstrate the launch vehicle guidance and control system during powered flight; achieve guidance'cutoff and evaluate system ac- curacy.

f. Demonstrate S-X/S-II dual plane separation and S-II/S-IVS separation.

g. Demonstrate launch'vehicle sequencing system.

h. Evaluate performance of the emergency detection system (FDS) in an open-loop configuration.

1' i. Demonstrate S-IVB stage restart capability.

j. Verify adequacy of the Command Module beat shield for xe- entry at lunar ixYcur* conditions.

3. Mission Profiles:

AS-501 will be launched from Launch Complex 39, Pad A, Kennedy Space Center (KSC); at a launch azimuth of 90"E of N. Shortly after liftoff (approximately 12 set) the vehicle begins a roll maneuver to at- tain a flight azimuth of 72% of N and maintains a near zero-lift (gravity turn) trajectory through the maximm dynamic pressure region. Afrer S-K burn and separation; S-II burn and separation, the first burn of rhe S-I"% will propel the S-IVB/IU/Spacecraft into a lOO-nautical- mile parking orbit using the Iterative Guidance Made (IGM). The vehicle will remain in this orbit for approximately two revolutions with its' longitudinal axis in the orbital plane and parallel to the local horizon. During the second revolution, when the vehicle is within tracking range of KSC, the S-ITS engine will be re-started to boost the vehicle into an elliptical atmosphere - intersecting waiting orbit with an apogee of approximately 9,000 nautical miles. Spacecraft separation occurs ap- proximately 590 seconds after injection into waiting orbit. The coast time in waiting orbit between S-ITS cutoff and Command Module (CM) re- entry is approximately 4.6 hours. Shortly after spacecraft separarion a service propulsion system (SPs) burn and navigational corrections will be performed to achieve lunar retnrn velocity and the proper re- entry corridor. Following the second SPS burn, Cormand Module/Service Module (CMM/SM) separation will occur and the Wwill be reoriented for a guided lifting reentry which will produce the heat load desired to test the CM heat shield at lunar returning velocity. Splash-down will be near Hawaii.

csn in

B...t to Earth &bit-S-K, s-ma B..d to S-II $

2

LIST OF FIGURES

Figure Title -

GENERAL

AS-501 Space Vehicle...... 6

Launch Complex - 39. ... , ...... 7

LC-39 Pad “A” Configuration ...... 8

SPACE/LAUNCH VEHICLE

4 Secure Range Safety System ...... II

5 Emergency Detection System (EDS) ...... 13

6 S-K/S-II Stage Flight Sequencing ...... 14

I S-II/S-1VR Stage Flight Sequencing ...... , . 15

3 S-IVB Stage Flight Sequencing ...... 16

9 Trajectory Information (Boost Phase) ...... 17

10 Guidance and Control System Block Diagram . . r 21

11 Vehicle Tracking Systems ...... 23

12 Space Vehicle Weight YS Flight Time ...... 25

S-IC STAGE

13 S-IC Stage Configuration ...... 27

14 F-I Engine system...... 29

15 s-x stage Propellant system ...... 31

16 S-IC Stage Tk.mst Vector Control System ...... 33

1T S-IC Stage Measuring System ...... 34

18 S-IC Stage Telemetry System ...... 35

19 S-IC Stage Electrical Power and Distribution System . . 36

S-II STAGE

20 S-II Stage Configuration ...... 39

21 J-2 Engine System - S-II Stage ...... 41

22 S-II stage Propellant System ...... 43

23 S-II Stage Propellant Management System ...... 45

24 S-II Stage Thrust Vector Control System ...... 47

25 S-II Stage Measuring System ...... 48

26 S-II Stage Telemetry System ...... 49

27 S-II Stage Electrical Power and Distribution System ...... 50

4 LIST OF FIGURES Figure Title m S-IVB STAGE 28 S-IVB Stage Configuration ...... 53 29 J-Z Engine System S-IVB Stage ...... 55

30 S-IVB Stage Propellant System...... 57 31 S-IVB Stage Propellant Management System...... 59

32 S-IVB Stage Thrust Vector Control System ...... 61 33 Auxiliary Propulsion System ...... 63

34 S-IVB Stage Measuring System...... 64

35 S-IVB Stage Telemetry System...... 65

36 S-IVB Stage Electrical Power and Distribution System ...... 66

INSTRUMENT UNIT

37 Instrument Unit Configuration ...... 69

36 Instrument Unit Measuring System...... 70 39 Instrument Unit Telemetry System...... 71

40 Instrument Unit Electrical Power and Distribution System . ~ . . . . 12

41 W/S-IVB Environmental Control Syst& ...... 75 SPACECRAFT 42 Spacecraft 017 Configuration...... 11

43 SC 011 Telemetry, USB & lJpaata systems...... 76 44 Spacecraft Eleotrical Power 2nd Distribution System...... 79

5 A-

S-U 5+a9a (NAA]

- 396 ‘3

-- A- A-

-396"

s-IC S+.9c(B.ein9)

6 H-P Gas Facility .^^ I \

Cr*wlerwny I

/I riquroz 7 rlobi,. *cr”/c. Strucfun

LsAnch Umbilical Tower

Pad -.--~-~__ F,,.. - ..__.

n The secure range safety systems on the S-K, S-II and S-I"B stages provide a communications link to transmit coded comands from ground starions to the vehicle during powered flight, providing a positive means of terminating the flight of an erraric vehicle by initiafing emergency engine cutoff and if necessary, propellant dis- persion.

Each powered stage contains tro UHF radio receivers. Both com- mand receivers on each of the thrre stages respond fo the same com,and signals, each providing a backup sysrem for the other.

The safety and arming device located on each stage is armed by a signal from the blockhouse before v;thicle ignition. Pollowing S-IVB cutoff the S-I"B range safety system is "safed" by a cornand from Range Safety Control to preclude accidental dertruct.

10 Chrg. I” ‘OX ranti ~.,rr.r.rr.r..,...,.rr.....r,,r,r,,,r., ,,,,,,, u ,,,,,I w,z ?zmeGmNcY DETECTlON SYSTEM (EDS)

The purpose of the EDS is to sense onboard emergency situations which arise during the boost phase of the fltght. On AS-501 the EDS will be flown in an open loop configuration which precludes autamatic abort.

The EDS is comprised of sensors which detect malfunctions and logic circuitry which initiate spacecraft displays and, in two cases, automatic abort of the M. With the exception of the g-ball, mounted on top of the LET, the EDS sensors are located in the launch vehicle. The system's relay logic is located primarily in the IU EDS Distribu- tor and the CM MissLo* Events Sequence Controller.

The EDS has two modes of operation; 'hamal", which generates abort cues and "automatic" which inifiates firing of the LES, and CM separation in the case of tw S-IC engines out or angular overrates during S-K powered flight. Figure is simplified black diagram of the AS-501 EDS. The automafic abort initiating portion of the system consists of ehe launch vehicle's rate sensing subsystem, the stage thrust sensing subsystem and the signal distribution and processing hardware which services these devices.

The angular overrate sensors (3 per axis in pitch, yaw, and roll) will initiate automatic abort of the CM during the period they are enabled (liftoff to about 136 seconds) whenever two sensors in any one axis simultaneously indicate excessive rates. Detection qf the over- rates is made by the se~-mr switch circuirry of-the Control Signal Processor in the I". The settings for these angular rate detectors are 5 degrees per second in pitch and yaw and 20 degrees per second in roll. The majority voting of the three switch outputs in each axis is done by relay logic in the FBS Distributor. A valid excess rate deci- sion is forwarded by the EDS Distributor to the CM Mission Events Se- quence Controller for abort initiation.

The S-IC stage engine thrust OK sensors (three per engine on all five engines) will also initiate abort during the period they are en- abled (liftoff to about 135 seconds) when the voted output of the sen- sors from any two engines indicates that the thrust of those engines is below the 89% level. These sensors monitor the F-l engine’s fuel inlet manifold pressure. Majority voting of the three sensors for each engine is done in the EDS Distributor. A valid two engines ant decision is sent to the Mission dents Sequence Controller for CM abort initiation. each H-i engine

Figure 5 Li

0 dutohmt’ic. Abort Capabilifq (Not active ou $A-501) t. 5-W Two ehlqiwes out (134.2+134..4) f!. Excessive aNqular rstes(13+.6+L34.a) I $ $-IC INboard Ewir-w Cutoff - 135 I A S-SC Outhard Ewqiwe Cutoff - 152 I ia s-11 ullaqe Triqqer-153 I A S-fC / s-11 SeParatiON-153 I rote : I &par. tiu\es 1 A s-11 mqiNe start-154 showu ate iN seconds wa- swed f row I U A Activate P. U. system -159 disco~r4ect I

A Water CoolaNt Vdlve OWN-160 I

r----l *2* 3 * Si of start Of smrt Pf Time Bose 1 The Base 2 Time BQSP 3

The BaSe DiViSiONS

Rqure 6

14 I A Separa+io? S-II SeCONd PlaNe--L83

A JettiSON l)?~--,s, i

A Cutoff S-it E&i!qiues--519 Aullaqe ~qwi+!L-~~~ I, A S-II/S-IV0 !L.epamtiou-520 I [ A IqNite S-IV B ENqi+m-523 I A~Acti~a+e’P~.SvsteM-528

$ CutQff S-IVB ENqiNe-665

p S-IVB UIlaqeENqiws ON-665

1 A Activate P.U. System-665 late : I I 4pprox. tiMe5 1 ihoww are i*l jecONd5 wasuped

-3-+-4-L-----5‘ start Of Start of Time Bose 4 TiMe Bose 5

Tiqure 1 I I I A S-IVB 1l.I.~ orJcl Lox ChilldowN i’u~ps .!w - 10,913 I I A S-IVB Restart Preparations-11,235

s-IVB Ulkqe Ew+les ON; - 11.235 b ~AS-IVBLH~ar4Lox ved valvesCIOSS -11,235 A~s-;m.y2 a,; Lox chiildoWN puyp off . I I I A S-IVB Ullaqo ENqines off -11,565 A 24 bp!rtiow S-IVB E~qi~e-11.570

A Acttivo{e I? U. system -11.515 I

I A S-IVB Tliqht Co&ml k&e: computer BUP~ Rppro%.tir4es Mocle off -11.890 itmWN are iN I 1ec0tds Mea- iured +ro~TlJ iiSCONHect.

---S---M 6 W 7- stall Of Start af titm? bare 6 tit.%?bo5a 7

Fiqure 6 i I 16 s-E73 24 Burn-- Start- 3hr, llmin., 44s Cutoff Q 3 hr., 17min.,~7s

hrkinq Orbit, &an Altitude-- 185 Km

Waitinq Orbit i Apogee-- 4.9666 km. I I , Cutoff Arminp .s I , IECO-- Switch S&&r command i i Plus sufficient downran9a 1 Tilt .Arrcst.- 1 velocity. ! OECO-- 145’.ls; Outbd. ‘~“9. propellant depletion cutoff enable.

ECO- - 483.9~ LOX depletion sensor-cutoff enable. LHz dcoletion sensor-cutoff

&unch Data Launch C.mp:e#.- 39 Pad A Launch Azimuth-90’5 of N Flight Arlmuth-7Z0 Eof N GU1DANCE AND CONTROL SYSTEM (G&C)

Function and De*cription

The G&C system provides these basic functions during flight: (1) stable positioning of the vehicle to the command position with a minimum amount of sloshing and bending, (2) a first stage tilt program which gives a near zero lift trajectory through the atmosphere and pro- duces reasonable end conditions at Outboard Engine Cutoff (OECO), (3) reduction of wind toads during the high dynamic pressure region, (4) steering commands during S-ID burn which guide the vehicle to a predetermined set of end conditions while maintaining a minimum propel- lant crajectary, (5) final cutoff signal, (6) attitude signals for the sensing, computing, and actuation elements of the G&C system. A block diagram of the hardware used to implement these functions is shown in Figure .

The Stabilized Platform (D-124&%) is a three gimbal configuration with gas bearing gyros and accelerometers mounted on the stable element. Vehicle accelerations and rotations are sensed relative to this stable element. Gimbal angles are measured by redundant resolvers and iner- tial velocity is obtained from accelerometer head rotation in the form of encoder outputs (also redundant).

The Launch Vehicle Data Adapter (LVDA) is an input-output device for the LVDC. These two components are digital devices which aperate in conjunction to carry out the flight program. This program performs the following functions: (1) P recesses the inputs from the ST-124M, (2) performs navigation calculations, (3) provides first stage tilt program, (4) calculates IGH steering commands, (5) resolves gimbal angles and steering commands into the vehicle system for attitude er- ror commands, (6) issues cutoff and sequencing signals.

The Control/EDS Rate Gyro Package contains 9 gyros (triplex re- dundant in 3 axes). Their outputs go to the Control Signal Processor (CSP) where they are voted and sent to the Control Computer for damp- ing vehicle angular motion.

The Control Computer sends commands to the S-IC, S-II and S-IVB engine actuators and to the Control Relay Packages based on signals from the LMA and rate gyros. These signals are filtered and scaled (see Figure IO), then smmed in magnetic amplifiers. 'I:& computer provides redundant operation during S-WB burn and coast.

The Control Relay Packages accept Conrral Camp"'-er commands and relay these comands to operate propellant valves in the Auxiliary Propulsion System CAPS). All relays and valves are redundant.

The 8 hydraulic actuators of the S-K and S-II stages are used to gimbal the outboard engines and provide control in all axes. The two hydraulic actuators of the S-IVB stage are used to gimbal the J-2 engine and provide control in pitch and yaw axes. The APS is used for control in all axes during S-ZVB coast. The Switch Selectors are used to relay sequencing commands from the LWA to other locations in the vehicle. The cutoff signal and time based events are issued through the Switch Se1cctors.

The vehicle is erected on the launch pad with position I at a 90'E of N azimuth. The Stabilized Platform is aligned to 72' azimuth durFng countdown and held in an earth fixed positioti, perpendicular to the gravity vector. A roll presetting of 18' is used to eliminate the at- titude error which would result from rhis difference in azimuth. The LWC operates in ground routines prior to GRR and attitude error signals are set to null. At the instant of cm, the platform becomes space fixed to establish the guidance coordinate sysrem and the LWC entei~ the flight mode. 1x1 this mode accelermeter processing, steering, navi- gation, telemetry, and other functions are performed exactly as rh6y are after liftaff. The Liftoff signal (IU lrmbilical disconnect) initiates time base 1. Eleven seconds later, the tilt program starts and the initial roll pre- setting is reset to zero. The vehicle rolls into alignment with the platform at l'/second. There is no active parh guidame during S-IC burn. Control is maintained by ginballing the four outboard engines on command from the Control Computer. Sequencing signals are issned to the various switch selectors to perform time dependent functions.

19 DIGITAL COMMAND SYSTEM CAPABILITY:

The fallowing summary describes the AS-501 Digital Command Systems' overall camand capability:

Function Periods of Acceptance

Inhibit FX~ TV + 100 SECO~S umii T6 and from T, + 10 seconds until end of life

Time base update Change the time to From Tg + 100 seconds until start coast phase T6 and from T7 + 10 seconds attitude maneuver until end of life

Tine base update Time base time is Pram Tg + 100 seconds until advanced or re- Tg and from T, + 10 seconds tarded at the nexf till end of life telemetry loss

Navigation updatt Navigation quanti- Frw Ti + 100 seconds until ties are reset at T6 and from T-/ + 10 seconds the time specified till end of life

Generalized switch Specified switch From Tz, C 115 seconds until selector selector function is T6 + 317 seconds and from issued at the first T7 + 10 seconds till edd of opportunity life

Sector dump Content.3 of speci- From Tg + 100 seconds until fied memory sector T6 and from T7 + 10 seconds are telemerered till end of life

Telemeter single From Tg + 100 seconds until memory location Tg and from TT + 10 seconds till end of life

Terminate Stop DCS processing From Tg + 1.3 seconds until and reset for a new Tb; from TI, + 115 seconds command till Tc + 317 seconds and from T7 + 10 seconds till end of life

Use S-US to achieve From Tq + 1.3 seconds till orbit =I Inhibits C-Band From Tg + 100 seconds until transponder switch- Tg and from T, + 10 seconds ing until end of life ST- 124M Stabilized - Platfclrm

Control Actuators 7,,//,,l, Illll,rll,-fIf,,, s-n sraqc To C.nTrol Actuators pIIt- Engiwb 1,1,3f4 * ( P, V,b R)

-““““““““““““““““““““““““““z S-ICC stag* Propellant consmq?ion during S-IC Stage flight (approximately 152 seconds) is approximately 4,320,OOO pounds. Propellant consumption during S-II Stage flight (approximately 365 seconds) is.approximately 938,700 pounds and during S-IVB Stage flight, including first and second burns, :approximately 459 seSopds$'is approximately 232,700 pounds.

In evenf of one engine of the S-X Stage malfunctions and is cutoff during flight, the remaining engines will consume the propellant intended for the "dead" engine. Burning time of the stage would increase, and the overall vehicle perfomance loss would be minimized.

VEHICLE WEIGHT DATA (Approximate)

Total af S-IC ignition Total at liftoff Total at s-IC O.E.C.0 Total at S-I* ignition 1;415;700 Total at S-II E.C.0 463,200 Total at S-IVB first ignition 352,000 Total at S-IVB first E.C.0 277,000 Total at S-IVB 2nd ignition 273,900 Total at S-IL'S Bnd E.C.0 119,100 Total at S-US/SC separation 67,100 6,220.700 Ibs (Total \lekcle at IqNit;oN) 6.000900 < 6,121.300 lbs (Total VehLcle at Liftoff )@ set)

S-XC PropellaM Corrsumptioti-28,4oCJ Ib+/sec. 3,000,000 /

S-lt Staqe Outboard Euqira Cutoff e 152 Sec.) Separatiohl (-153 Sec.) S-II Staqe 1qNitiot.I (~154 Sec.) (1,415.700 \bs)

1.000.000 S-II PmpellaNt CohlsUMptoN - 2,590 Ibs/sec. Vehicle Weiqht ( FwNd5) S-II Staqe LqiNe Cutoff (-523Sec.) Staqe SePartiiON (-520 Sec.) 4oo,ooa

100.00~

30,ooc

Space Vehicle Weiqht vs lliqht TiM e

-Fi& The 8-X stage is approximately 138 feet long and 33 feet in di- ameter and has five liquid-fueled Rocketdyne P-l engines which generate a total thrust of 7,500,OOO pounds. The engines are supplied fuel by a bi-propellant system of liquid oxygen (LOX) as the oxidizer and W-1 as the fuel.

The S-TIC stage structure consists of a thrust structure to which the engines attach, an W-1 fuel tank, a LOX tank, an interrank SYNC- ture separating the LOX and fuel ranks, and a forward skirt structure which provides an interface surface for the Saturn g-11 stage.

26

F-l ENGINE OPERATION

The F-l Engine is started by ground support equipment. The start signal ignites pyrotechnic igniters in the gas generator which permits LOX under tank pre*sure to discharge into the thrust chamber. When the LOX valve is partly open RP-1 and LOX under tank pressure flows to the gas generator combustion chamber accelerating the turbopump increasing the LOX and RF-1 discharge pressure. When the W-1 discharge pressure reaches approximately 375 psig a valve in the bypergol cartridge opens allowing LOX and W-1 to build up pressure against the hypergol burst diaphragm. At approximately 500 psig the diaphragm will rupture allowing hypergol and RF-1 to enfer the thrus't chamber causing spontaneous combustion upon contact with the LOX, thereby establishing primary ignition. As thrust pressure builds up the U-1 valves open admitting W-1 to the thrust chamber and the transition to mainstage operation.

Inboard engine is cutoff by a signal from the I". Outboard engines are cutoff by optical type LOX depletion sensors with fuel depleting sensors as backup. A command from the IU supplies a coormand to the switch selector to enable the outboard engine cutoff circuitry. When two or more of the four LOX level sensors are energized, a fimer is activated. Expiration of the timer energizes the stop solenoid far each engine which energizes prevalve close relays. Closing of the prevalves inrerruprs propellant flow and terminates engine operation. RP-I LOX TIP-1 i 1 I I

1,500.000 I bs. Thrust

fiaure 14 The S-E Stage propellant system is composed of one LOX tank, one RF-1 tank, prope11anc lines, control valves, uents, and pressurization subsystems. Loading of LOX and W-1 tanks ;s controlled by ground computers. W-1 loading is compieted at a considerable time prior to start of LOX loading. LOX bubbling begins and continues throu@, the LOX tank loading to prevent possible LOX geysering. Approximately 90 seconds prior to ignition command rhe RF-: tank is pressurized from a ground source. Approximately 60 seconds prior to ignition command the LOX tank is pressurized from a ground source. Prior to start of automatic sequence and up to 72 seconds before liftoff ground source helium is bubbled through the LOX lines and tank to prevent stratification in engim LOX suction lines. After liftoff fhe LCX tank pressurization is maintained by GOX converted from LOX in the heat exchanger. The W-1 tank is pressurized with He stored in bottles in the LOX tank and heated by passing the He through the heat exchanger.

100 --+ I ’ 80 I ‘. I I” L 60 IqNitiON commaNc 51 D 4a

I reTmuk - w-1 jtart Sy5tet.7 .oadiNq RD.1

Low Re - Pressurize TscuK-

30 31 cu. +t. 3000 p&9. He Bottle5 (4) for W-1 Tamk Pressuriz Veut C Relief Valve Oprlv -25.5 prig Close -24.0 psiq

17 iwdiau\eter -A- Lox Fill L Drain 33rierqewy DraN I Ved &Reiiet Valve f- Ye& -Relief 30.5 psia 35.0 wia

I -RP-1 ?Gll CDrair 12 iw. diw LiNe er etupe) iN hNeS 01 Liftoff - b‘tS0 Ibs.

Total propellant at liiioif 4,3%OOO bs. Total pmpel\ard co~suwwd dter l&t& -4,320,2OO IbS.

1 ! /f ‘vehicle Thrust S*ructure.A+tuator AitachueM Poinrt Hydraulic Actustor (2 per .eN9iW) Tw’L,~~~~,, I I fuel dlscharqe \ Servo vcllve C‘JNtt’OiS duct cktuator ~~OV.SMPH+

Actuators III

CaNted go NOMINAL thrust ~---. -- ---Signal SOUtXCs

fhcrnocoupls

Measuring Rack 2, in S-K stage \

r- Sign*ls (liF:off,c.o., OtCJ F

14 Antennas

Divider

L-rJ Multicoupler Multicouplrr ‘r Data InDuts I

7 r&-o-rocket firing Until IOSS of signal.

Data _ Inputs Located in -. 231.9 Ins Thrust Frame Aru P-l PCMjDDA4 4 = ,qss.v -) Hardwin ToGSE

Figure 18 -- Ij

35 Mraruring Lnstruma~tati~n d r-i Diatri butor TRtcmctry systems

Battery Battery No. 1 No.2 640 amp. min. 1250 amp. min.

6reund P.w.r Main Power Distributor 2.9 vdc

In~ttrumrnt.tion .$ -iolomdry systems

*E”ginr system (Thrust) . ED5 system

Nate: 1 All Components Show” ~tl~chsniral Systems Arr Located I. The Thrust . DDAS sy5tms Frame Area Except as N&cd. . Rang* safety systems - staging System - Squcnciny System

36 37 The stage sfrucfure includes: an aft interstage, an aft sk.irt and thrust structure, a heat shield, a LOX tank, an LH2 tank, and a fc. skirt.

The stage has five J-2 engines which generate a total thrust of 1,000,000 pouuds.

Two recoverable film cameras will be flown an the S-II stage of rhe first two Saturn V launch vehicles. 'These cameras are mounted an the thrust structue as shown on opposite page. The primary objective of these cameras is to view second-plane S-X/S-II separation. The secand- ary objectives are to view S-IC/S-II first-plane separation and J-2 rngine ignition,

The cameras are turned on by the switch selector shortly before first-plane separation and operate for approximately 40 seconds. The film is 'marked" one-tenth of a second before first-.plane separation, one-tenth of a second after engine start and one-tenth of a secclnd after second-plane separation. The camera capsules are ejected ap- proximately 8 seconds after second-plane separation.

Imediately following ejection, the camera capsule stabilization flaps are deployed. After the camera capsule descends to an altitude Of 4,3?0 meters, a paraballoan is inflated, which CRUSTS the stabili- zation flaps TO fall away. Six seconds after the paraballoan is in- flated, a recovery radio tranmit.te+ and flashing light beacon located an the paraballaon are turned on.

After touchdown, the camera capsule effuses a dye marker to aid visual sighting of the capsule, and a shark-xepellant co protect the camera capsule, paraballoon, and the recovery team.

38 -2 Fill # Drain LH2 F..dl;.c

Ullrg.- Aockot. (8)

Lights (28) For Cameras

--5tag. wcigy . Dry :- 88,200 Lbs. . fit s-IC Ignition:-- 1,035,300Lbs. . A+ S-II Cutof+:- 103,500 Lbs. The operating cycle of the 5-2 Engine ccmsists of prestart, starf, steady-state operation and cutoff sequeoces. During prestart, Lox and LIZ2 flow through the engine to temperature-condition the engine components, and LO assure the presence of propellane in the turbopumps for starting. Following a timed cooldown period, the start signal is received by rhe sequence controller which energizes various control solenoid valves to open the propellant valves in the proper sequence. The sequence controller also energizes spark plugs in the gas generator and thrmt chamber fo ignite the propellant. In addition, chr sequence concro11er releases P,Wz from the start tank. Abe GHz provides the initial drive for the turbo- pumps that deliver propellant fo the gas generator and the engine. 'i&o propellant ignites, gas generator output accelerates the turbopumps, and engine Chrust increases to main stage operation. At this rime, the spark plugs are de-energized and the engine is in steady-state operation.

Steady-state operation is mintained until a cutoff signal is received by the sequence controller. The sequence controller de-energizes the solenoid valves which in turn close the engine propellant valves in the proper sequence. *s a result, engine Chrust decays and the cutoff sequence is compleLe. Lsr Propel!ant utilization valve

I-!- GHe for LHz tank pressurizatiob 4 JI

1He IIE -:*cll

200.00d Ibs. Thrust

tic&Ire 21 The S-II Stage prope:lant system is composed of integral LOX/LH~ tanks, propellant lines, control valves, vents, and prepressurizaCion subsystems. Loading of propellant tanks and flow of propellants is controlled by the propellant utilization systems. The LOX/LH2 tanks are prepressurized by ground source gaseous helium. During powered flighr of the S-II Stage, the LOX tank is pressurized by GOX bleed from the LOX heat exchanger. The LH2 tank is pressurized by OH2 bleed from the chrusc chamber hydrogen injector manifold: pressurization is maintained by the LH2 Pressure Regulator.

S-II PKOPELL4NT LOAD AND OPERATIONAL SEQUENCE

,------1 --* ‘- mouqal/P~ \ 1000 qal/he4 96% , \ 98% Start Autowotic L!!z-j I I I

I - I I + 7.hr. -6.5hr. -S-T&. -4.shr. -T.t.w toLi:toff -1e7 SK. -6 ie c.

A I Start LoadiNq LHZ LH2 Tat& -153.000 Ibs. dt IqNitiQN

Lam. TaNk -‘?92.6~ lb5. dt IqHltiQN

DraiN

HeatErchahlqer--.d...,,. F-+-Y I\ coverts Lox t Gox far Lox TaNk pressurizu&iv~ duriNq S-11 powered fliqht .

Total ~ropellcwt dt liftoFf -9%. .M !L-. Total propeik~t cousu~ed dfter lif*f -936x00 It6

FiquW 22 I[ The propellant utiiization (PU) system controls loading and engine mixture ratios (LOX to LH2) to ensure balanced consumption of WX and LH2.

Capacitance probes mounted in the LOX and LH2 containers mm&for t-be mass of the propellants during powered flight. At PU activaeiou (5.5 seconds after J-2 ignition) the capacitance probes sense the LOX to LA2 imbalance and commands the engine to bum at the high rate engine mixture ratio of 5.5:1. When the high mixture ratio is removed, the PU system will then command fhe engine to burn the reference mixture ratio of 4.7:1, striving for siw.ultaoeous depletion of LOX and LHz for maximum stage performance. Engine cutoff is initiated when any two of the five capacitance probes in either Cask indicate dry. Telemetered I Xrror Signal 4

Telemetetsd V4lvr Command Siqnal

TO other 4 enqiqes F’

to Ground E Loading computer 4 Telemetry

CONtlOlValve Teietietered Valve P&t iou SiqWal

Tiqure 23

,’ 4 To Blockhouse thru pwflighf Total6 + 79 * data acquisitioe System t+ All ESE Display F; I;

r I

Microphone _ (acoustic 1

48

DistribUtDr

------_--__ - - .&&--_ 5kirt Thrust Framr A,%&

Proprllmt Monitoring FMahagrmenf Systems

POW.%V Range Safety Recirculation system* I Batterv No. 1 4

--Figure t27 Intentionally Left Blank

57 ,,,, 52 --++t

lk Top View Looking Aft

T&mr+qAntcnnrs (4) Forwwdskirt A

Auxiliary Tunnel

comm*n BUlKhDId

S-D?3 sta3* Weight * .Dry:-~6,500 Lb.. *At Ignition:~Zbz,ZoO Lbs. ‘Al- I& Cutoff:-- lE7,3OOLbs. *At 2Ud&Cutoff:--29,300Lbo. SEx*cludrs Wt. of aft.. interstage Thz operating cycle of the J-2 Engine consists of prestart, start, steady-state operation and cutoff sequences. During prestar?, LOX and LH2 flow through the engine to temperature-canditian the engine components, and to assure the presence of propellant in the turbopumps for starting. Following a timed cooldo-m period, the start signal is received by the sequence controller which energizes various control solenoid valves to open the propellant valves in the proper sequence. The sequence controller also energizes spark plugs in the gas generator and thrust chamber to ignite ctle propellant. In addition, the sequer.ce controller releases ai* from the start tank. The GE2 provides the initial drive for the turbo- pumps that deliver propellant fo rhe gas generator and the engine. The propellant ignites, gas generator output accelerates the turbopumps, and engine thrust increases to main stage operation. At this time, the spark plugs are de-energized and the engine is in steady-state operation.

Steady-state operation is maintained until a cutoff signal is received by the sequence controller. The sequence controller de-energizes the solenoid valves which in turn close the engine propellant valves in the proper sequence. As a result, en.ci.ine thrust decays and the cutoff sequence is complete.

I

.

COMMAND N TlME FROM IGLIITIOH

54 GHz for Ltiz tank p~essUrizC,tiOC, i !

GCJS for LOX ton* pressurization A

1 200,000 Ibs.Thrust

II”.’

fiqure 29 -- The s-IW stage propellant system is composed of integral LOX/L~ tanks, propellant lines, control valves, YeDts, and pressurization subsystems. Loading of the propellant tanks and flaw of propellants is controlled by the prope1Lanr utilization system. Both propellant tanks are initially pressurized by ground solnrce cold helium. LOX tank pressurization during S-WE *tage bum is maintained by helium supplied from spheres in the I,Ez tank, whLch is expanded by passing through the helium heater, to maintain positive pressure across the common Lank bulkhead and to satisfy engine net positive suction head. The LHz pressurization strengthens the stage in addition to satisfying net positive suction head requirements. After engine ignition the presst,re is maintained by GH2 tapped from the engine supply 1------L---- start AUtOttlstiC SequenceT-l liftoff t A I Ignition9 tom owl

I I I

15 ht. -7.0 hr. -4.i hr. -319 hr. I -f;me to liitotf i 3 5 cu.+t. 31W sia rjHs ~sbhores(8). IN. PIiqht Los rqqrrk preswruut&w ,

i ; r-' LHITciHk I -41.200 lb5 al lqrri+ioN / i I ‘Ml I 1\r 1 j [I- l ’ I// I I I ,c /f--Los Fill 4 Drsiu ’ I-I-_ lL ,r;KjLI- Hpot ExchaNqef u

Total propeitnb&t at i;ftoff -234.100 4trs. TO-M propellaN+ com%med after liftoFf-232.7m lbs.