atlas-agena performance for the 1967 mariner venus mission 19690020216_1969020216

8/3/2019 Atlas-Agena Performance for the 1967 Mariner Venus Mission 19690020216_1969020216

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NASA TM X-1826

ATLAS-AGENA PERFORMANCE FO R THE 1967 MARINER VENUS MISSION

By L e w is R e s e a r c h C e n t e r StaffL e w is R e s e a r c h C e n t e r

Cleveland, Ohio

NATIONAL AERONAUTICS AND SPACE ADMINISTRATIONFor sale by the Clearinghouse for Fed era l Scientifi c and Tec hni cal Informotion

Springfield, Virginia 22151 - CFSTI price $3.00


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ABSTRACTThe Atlas-Agena successfully launched a Mariner unmanned deep space probe

during 1967 to fly by the planet Venus. This report discusses the performance of theAtlas-Agena from lift-off through the Agena posigrade maneuver. The objective of themission w as to extend scientific knowledge about the size and mass of Venus and thedensity and composition of the planet's atmosphere.

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CONTENTSPage

1.SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1I1. INTRODUCTION by Roy K. Hackbarth . . . . . . . . . . . . . . . . . . . . . . 3

I11. LAUNCH VEHICLE DESCRIPTION by Roy K. Hackbarth andEugene E. Coffey . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5Tv . TRAJECTORY AND PERFORMANCE by James C. Stoll and

KennethA. Adams . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11TRAJECTORY P L A N . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11TRAJECTORY RESULTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

Lift-off Through Atlas Booster Phase . . . . . . . . . . . . . . . . . . . . 12Atlas Sustainer Phase . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13Atlas Vernier Phase . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13Agena Engine Firs t-Burn Phase . . . . . . . . . . . . . . . . . . . . . . . 14Agena Engine Second-Burn Phase . . . . . . . . . . . . . . . . . . . . . . 14Post-Second-Burn Phase . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

V. ATLAS VEHICLE SYSTEM PERFORMANCE . . . . . . . . . . . . . . . . . . . 23VEHICLE STRUCTURE SYSTEM by Richard T. Barrett . . . . . . . . . . . 23

Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23

25Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

29Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

HYDRAULIC SYSTEM by Eugene J . Cieslewicz . . . . . . . . . . . . . . . . 32Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32

PNEUMATIC SYSTEM by Eugene J . Fourney . . . . . . . . . . . . . . . . . 35Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36J a m e s L. Swavely . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40

PROPULSION SYSTEM by Charles H. Kerrigan . . . . . . . . . . . . . . . .PROPELLANT UTILIZATION SYSTEM by Clifford H.Arth . . . . . . . . . .

GUIDANCE AND FLIGHT CONTROL SYSTEM by Dean W . Bitler and

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ELECTRICAL SYSTEM by Clifford H. Arth . . . . . . . . . . . . . . . . . . 49Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 9Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

TELEMETRY SYSTEM by Edwin S. Jeris . . . . . . . . . . . . . . . . . . .Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

FLIGHT TERMINATION SYSTEM by Edwin S. Jeris . . . . . . . . . . . . .Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

V I. AGENA VEHICLE SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . .VEHICLE STRUCTURE SYSTEM by Robert N . Reinberger . . . . . . . . . .

Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .PROPULSION SYSTEM by Robert J. Schroeder . . . . . . . . . . . . . . . .Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Historical Discussion of Engine Anomaly . . . . . . . . . . . . . . . . . .Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Subsystem Performance . . . . . . . . . . . . . . . . . . . . . . . . . . .

SHROUD-ADAPTER SYSTEM by C. Robert Finkelstein . . . . . . . . . . .

GUIDANCE ANI) FLIGHT CONTROL SYSTEM by Howard D. Jackson . . . .ELECTRICAL SYSTEM by Edwin R. Procasky . . . . . . . . . . . . . . . .COMMUNICATION AND CONTROL SYSTEM by Richard L . Greene . . . . .

VII. LAUNCH OPERATIONS by Frank E. Gue and Alvin C. Hahn . . . . . . . . . .PRELAUNCH ACTIVITIES . . . . . . . . . . . . . . . . . . . . . . . . . . .COUNTDOWN AND LAUNCH . . . . . . . . . . . . . . . . . . . . . . . . . .

VIII. CONCLUDING REMARKS. . . . . . . . . . . . . . . . . . . . . . . . . . . . .APPENDlXES

A . EQUENCE OF MAJOR FLIGHT EVENTS. MARINER VENUS 6 7by Richard L. Greene . . . . . . . . . . . . . . . . . . . . . . . . . . . .

495 15 15 1555 55 55 75 75 75 7595 96 06 26 26 36 5686 86 9737 37 375757677777779

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B . AUNCH VEHICLE INSTRUMENTATION SUMMARY. MARINER VENUS 67b y Edvvin S. Jeris a n d R i c h a r d L. G r e e n e . . . . . . . . . . . . . . . . . . 8 2b y R i c h a r d L. G r e e n e . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 9b y R o b e r t W . Yo rk . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9 2

C . RACK ING AND DAT A ACQUISITION. MARINER VENUS 67D . EHICLE FLIGH T DYNAMICS. MARINER VENUS 67

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I. SUMMARYThis report evaluates the flight performance of the Atlas-Agena launch vehicle insupport of the Mariner 67 flyby of the planet Venus. The Atlas-Agena launch accurately

placed the Mar iner into a Venus transfer ellipse and the Mariner passed within 3700 kilo-me te rs of the planet Venus on October 19, 1967.

The Mariner 67 space vehicle was successfully launched fro m the Ea ste rn TestRange on June 14, 1967. The Atlas boosted the Agena-Mariner into the proper sub-orb ita l coast e llipse and separation of the Agena-Mariner fro m the Atlas was successful.The Agena first burn placed the Agena-Mariner into the des ired parking orbit . After anorbita l coast period, the Agena second burn placed the Agena-Mariner into the desiredVenus transfer ellipse. The Mariner was separate d from the Agena. The Agena thenperformed the required posigrade maneuver to ensure the Agena would not in ter fere withMariner o r impact the planet Venus.

A l l Atlas sys tem s performed satisfactorily. The Mod III radio guidance sys tem per-formed satisfactori ly although a backup method w a s used by the ground station to acquirethe vehicle. All Agena system s performed satisfac torily except the Agena engine expe-rienced chamber press ure variations during second burn. These press ure variationshad no effect on the attainment of the mission objective.


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11. INTRODUCTIONby R oy K. Hackba r t h

The purpose of the Mariner Venus 67 mission was to obtain scientif ic data on thephysical c haracter is tic s of the planet Venus and the composition of the Venusian atmo-sphere . The objective of the launch vehicle was to inject the Mariner into a prescribedVenus tran sfe r trajectory and, after spacecraft separation, to perform a posigrade ma-neuver to ensure the Agena would not interfere with the Mariner or impact Venus.

This wa s the fifth Atlas-Agena launch in support of Mariner flybys of Mars and Venusand the thi rd launched under the direction of the Lewis Research Center.

Mariners I and I1 we re launched toward Venus in 1962. Mariner I was unsuccessfuldue to a booster malfunction and Mariner I1 performed the first successful flyby of Venusin December 1962. Mariners I11 and IV were launched towards M a r s in November 1964.Mariner I11 was unsuccessful due to failu re of the nose fair ing to se parate and Mariner IVperformed a successfu l flyby of M a r s in July 1965. Mariner V (Mariner Venus 67) waslaunched toward Venus on June 14, 1967.

This report evaluates the flight performance of the Atlas-Agena launch vehicle forMariner f rom lift-off through the Agena posigrade maneuver.

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111. LAUNCH VEHICLE DESC RIPTIONby R o y K. H a c k b a r t h a n d E u g e ne E. Cof fey

The Atlas-Agena is a two-stage launch vehicle consisting of an Atlas first stage andan Agena second s tage connected by a booster adapter. The composite vehicle (fig. 111-1)including the spacecraft shroud and booster adapter is 3 1 . 7 meters (104 f t ) in length.The vehicle weight at lift-off is approximately 125 600 kilograms (277 000 lb). Figure111-2 shows the Atlas-Agena launch vehicle lifting off with the Mariner Venus 67 space-craft.

The A t l a s SLV- 3 (fig. 111-3) is 2 1 . 3 4 meters (70 f t ) long and is 3 . 0 5 meters (10 f t ) indiameter except for the forward section of the tank which is conical and taper s to a diam-eter of about 2 meters (6 f t ) . Atlas is propelled by a standard Rocketdyne MA-5 propul-sion sys tem consisting of a booster engine having two thrust chambers with a total thrustat sea level of 1 4 6 7 . 9 ~ 1 0 newtons (330 000 lb); a sustainer engine with a thrust at sealevel of 2 5 3 . 5 5 ~ 1 0 newtons (57 000 lb); and tw o vern ier engines, each with a thrust at3sea level of 2 . 9 8 ~ 1 0 newtons (669 lb). A l l engines use liquid oxygen and high-gradekerosene propellants and are ignited pr io r to lift-off. The booste r thrust chambers aregimbaled fo r pitch, yaw, and ro ll control during the booster phase of flight. This phaseis completed when the vehicle acceleration equals about 6 gs. The booster engines arejettisoned about 3 seconds after booster engine shutdown. The sustainer and vernier en-gines continue to burn fo r the susta iner phase of flight. During this phase, the sustainerengine is gimbaled for pitch and yaw control, and vernier engines are gimbaled for ro llcontrol. The sustainer engine burns until the vehicle achieves the desir ed suborbitalparameters as determined by the ground radio guidance system. A f t e r sustainer engineshutdown, the ver nier engines continue to burn for a short period pri or to the A t l a s -Agena separation. During thi s phase, the vern ier engines are gimbaled to provide vehi-cle attitude control and fine trajec tory corrections . After vernie r engine shutdown, theAtlas is sev ere d from the Agena by the firing of a Mild Detonating Fuse (MDF) systemlocated on the booste r adapter. The firing of a retrorocket system, mounted on thebooster adapter, then separate s the A t l a s booster adapter from the Agena.

The second-stage Agena and the shroud protecting the spacecraft are shown in fig-ure 111-4. The diameter of the Agena is 1 . 5 2 meters (5 f t ) , and the length of the Agenashroud is about 10 meters (34 f t ) . Agena is powered by a model 8096 Bell Aerosystemsengine with a rated vacuum thrust of 7 1 . 1 7 ~ 1 0 newtons (16 000 lb) and has a two-burn

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capability. This engine uses unsymmetr ica l dimethylhydrazine and inhibited re d fumingnitric acid as propellants. During powered flight, pitch and yaw control are provided bygimbaling the Agena engine, and ro ll control is provided by a cold gas (mixture of nitro-gen and tetraf luoromethane) system. During per iods of nonpowered flight, pitch, yaw,and roll control are provided by the cold gas system. The cold gas system and a posi-grade rocket on the Agena are used to complete a posigrade maneuver after spacecraftrelease. A metal shroud is used to provide environmental protection for the spacecraftduring ascent. This shroud is jettisoned after Atlas vernier engine shutdown just priorto Atlas-Agena separation.

-- -Over-the-nose shroudgena station 85.65

Agena station 526 IAtlas station 502

1

Atlas station 1133

.--Agena booster adapterd-extension (stub)

Atlas stationFigure 111-1. - Space vehicle profile, Marin er Ven us 67.

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Figure 111-2. - Atlas-Agena vehicle l ifting off wi th M a r i n e r V e n u s 67 .

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Liquid-oxygentank- --Boostert h r u s t c h am b e r 2 -

In te rmedia tebulkhead--.Fuel tank-,

Fuel tankp r e s s u r i z a -t ion l ine-.--

Sta t ion,--Forward ta n k Liqu i d-oxygen4 bulk hea d boiloff valve -\

tIV

--.

,-Cable fairing- 45.0

-960.0-Liquid- oxygen line1- -1 pod

/-Umbilical panel-1133.04,-iquid-oxygenf i l l a n d d r a i n

Figure 111-3. - Atlas SLV-3 conf igura t ion .

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,-Retrorockets (a,i 180"apart

Oxygen f i l l door-, /;%uel f i l l doorNitrogen fill door-/Retrorocket fairing (2)

BoDster adapter s tructure- ,Roller access doors (8)-

,-Command de stru ct ch arq e\\\\

'7 -S e l fdes t ruc t charge f a i r ing, I,/ L s epara t ion r a i l s (4),,

LMild detonating fuseseparation ring\\\- Posigrade rodtet

CTank sect ion

Figure 111-4. - Agena conf igurat ion, Mariner Venus 67.

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IV . TRAJECTORY AND PERFORMANCEb y James C. Stoll and Ke nn e t h A. Adam s

TRAJECTORY PIANThe Atlas boosts the Agena-Mariner onto a prescr ibed suborbital coast ellipse, and

the Agena perfo rms two engine burns to place the Mariner onto a Venus tra nsfe r orbit.The Atlas flight consis ts of thr ee powered phases: a booster engine phase, a sustainerengine phase, and a vernier engine phase. Shroud ejection, followed by Agena separa-tion, occurs after the vernier phase is completed. Following Atlas-Agena separation,the Agena engine first burn places the Agena-Mariner onto a circul ar parking orbit at analtitude of 1 8 5 . 2 kilometers (100 n. mi .) . The Agena-Mariner coasts in the parking or-bit to a predetermined point where the Agena engine is reignited. The Agena enginesecond burn places the Agena-Mariner onto a Venus transf er orbit with the requi red en-ergy to intercept the planet. The Marin er is then separated f rom the Agena and theAgena perfo rms a posigrade maneuver to ensure that it will neither inte rfer e with thespacecraft nor impact Venus.

Lunar and planetary missions require a flight azimuth which is a function of lift-offtime. In practice, flight azimuths are precalculated for the midpoints of successive 15-minute intervals throughout the daily launch window. These intervals are identified as"Launch Plans. ' ' The precalculated flight azimuth for the designated launch plan is pro-grammed in the Atlas airbo rne flight control progra mmer. The progra mmer commandsthe vehicle to r ol l from the pad azimuth to the precalculated flight azimuth during the13 seconds of flight star ting at lift-off plus 2 seconds (T + 2). If launch occurs at a timeother than the instant corresponding to the programmed flight azimuth, a discrepancyexists between the programmed flight azimuth and the desi red flight azimuth. This er ro rrequires a guidance correct ion during ascent. The correction, generated by the groundradio guidance computer, is accomplished by yaw steering during the sustainer phase ofthe Atlas flight.

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TRAJECTORY RESULTSLift-off T h r o u g h A tl as Booster Phase

Marin er Venus 67 (MV-67) was successfully launched from Complex 12, EasternTest Range, on June 14, 1967 at 0101:OO. 1 76 Easte rn Standard Time using Launch Plan14C. This launch plan required a programmed flight azimuth of 102.3'. The launch padazimuth is 105.18'. A comparison of actual and expected times fo r the significant flightevents for this mission is presented in table IV-I and in appendix A.

Winds at launch wer e predominately fro m the eas t fro m the surfac e to about 3048meters (1 0 000 f t ) altitude. Beheen 3048 and 6096 meters (10 000 and 20 000 f t ) altitude,the winds wer e f rom the northeast and above 6096 meters (2 0 000 f t ) from the northwest.The winds were light and had only a minor effect on the vehicle flight path. Above 6096meters (2 0 000 f t) altitude, the winds were tail winds and tended to de pre ss the tra jectoryslightly. A maximum wind velocity of 1 8 . 3 meters pe r second (60 ft/sec) f ro m the north-east occurred at an alt itude of 1 4 6 7 4 . 9 meters (48 146 ft). Wind data are shown in fig-u re N - 1 . Abrupt changes in wind velocity at altitudes between 9144 and 15 240 meters(30 000 and 50 000 f t ) produced strong wind she ars .

The maximum vehicle bending response was calculated to be 4 3 . 9 percen t of thecrit ica l value at the Mariner-spacecraft adapter interface (Agena Station 247) and tooccur at an a ltitude of 9404 meters (30 853 ft). The maximum booster engine gimbalangle was calculated to be 4 9 . 1 percen t of the available gimbal angle in the pitch planeand to occur at an altitude of 9 4 7 5 . 3 meters (31 087 f t ) . The data used for these calcula-tions we re obtained from the T - 0 (lift-off) weather balloon.

Radar tracking data show that the vehicle flight path was slight ly lower than the ex-pected trajectory during the A t l a s booster phase of the flight. This deviation resultedprimarily from tail winds and an approximate 0.4 ' excess total pitchover command bythe flight control system during the booster phase. At booster engine cutoff (BECO), thevehicle position was about 1 . 9 8 kilometers ( 1 . 0 7 n. mi .) below and 1 . 3 4 kilometers(0.72 n. mi. ) downrange of the expected position.

The tra jec tory deviated in the horizontal plane only slightly fro m the expected tra-jectory, resulting in a position 0 . 0 6 1 kilometer ( 0 . 0 3 3 n. mi. ) left at BECO. This ef-fect was due primarily to a 0.24' exces s vehicle rol l during the programmed rol l ma-neuver.

During the booster phase of flight, the capability of the flight control system to ac-cept Mod III Radio Guidance commands was enabled at T + 80 seconds; however, pitchsteering commands could only be transm itted fro m the ground station between T + 100and T + 110 seconds. No booster pitch ste ering oc curred during this interval since thevelocity vector angle dispersion in the pitch plane was less than the predetermined

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threshold fo r booster steering. Radio guidance yaw s tee ring was not progra mmed to beused during the booster phase of flight. BECO occurred at T + 128.6 seconds by groundrad io guidance command at a vehicle longitudinal accelerat ion of 5.9 g's. The accelera-tion level at BECO was 0.1 g lower than expected but within the guidance tolerance of&O. 2 g. The booster engines were jettisoned at T + 131.6 seconds.

A t l a s S u s t a i n e r P h a s eThe trajectory remained depressed and left of the expected trajectory during the

sustain er phase. Tracking data indicate tha t at sustainer engine cutoff (SECO) the ve-hicle position was about 5.79 ki lom ete rs (3.13 n. mi .) lower than, l. 01 kilometers(0.55 n. m i. ) left of, and 0.73 kilometer (0.39 n. mi. ) downrange of the expected posi-tion.

mands caused the vehicle to pitch up approximately 5.7' and yaw left 2.4'. These ma-neuvers wer e made to compensate for the low tra jectory and to steer the vehicle to thedes ired flight azimuth. N o correc tions wer e made fo r the cr os s range displacementer r or s accumulated during the booster phase. Therefore, these er ro rs were in evidenceat SECO. SECO was commanded by ground radio guidance at T + 296.7 seconds,0.6 second earlier than expected.second (10 t/sec) higher than expected. The total Atlas performance including this ve-locity increment and the depr essed trajec tory of the vehicle was consistent with the de-si re d energy fo r the expected suborbital coast ellipse. The suborbital coast ellipse pa-rameters are given in table IV-11.

Sustainer steer ing was initiated at T + 138.4 seconds. The initia l steering com-

The vehicle velocity (re lat ive to the rotating ea rth ) at SECO w a s 3.05 meters per

A t l a s V e r n i e r P h a s eVernier engine thrust duration after SECO was approximately 21.2 seconds. During

the vernier phase pitchdown and yaw, right steering commands were issued by groundradio guidance in or der to place the vehicle in the prope r attitude before Atlas-Agenaseparation. These commands displaced the vehicle 0.5' right in yaw and 3.0' down inpitch. Vernier engine cutoff (VECO) occurred by ground radio guidance command atT + 317.9 seconds, 0.5 second later than expected. Atlas inser tion veloci ties at VECOare given in table N-III.ond later than expected. This delay was due to a guidance equation requ irement that the

Shroud separation was commanded by radio guidance at T + 320.1 seconds, 0.6 sec-

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time r tolerance. The engine second-burn duration (measured from 90 percent chamberpr es sur e to velocity meter cutoff) was 94.4 seconds, 1.2 seconds shorter than expected.During the second burn, a momentary decrease in chamber pre ss ure occurred. Thepre ssur e returned to a value slightly greater than the i a1value. The overall effect ofthe chamber pr ess ur e variation was to increase engine average thrust, hence, the short erburn duration. Velocity me te r cutoff indicated that the proper velocity had been gained.Thrus t decay velocity was 13.32 meters pe r second (43.7 ft/sec) as compared to a pre-dicted decay velocity of 15.12 meters p er second (49.6 ft/sec).

Post - Second - B u r n P h a s eThe spacecraft w a s separa ted from the Agena at T + 1576.7 seconds. The space-

craft trajectory parame ters at final injection a r e given in table IV-V. Target point pa-ram ete rs for the spacecraft a r e given in table IV-VI.firing was initiated at T + 1579.7 seconds and w a s successfully completed at 1892.5.Integration of the resul ting Agena position and velocity vec tors to Venus gave the missdistances shown in table IV-VII.

The Agena posigrade maneuver consisting of a yaw maneuver and a posigrade rocket

The Mariner w a s placed in its desire d orbit and performed satisfactorily.

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TABLE IV-I. - SIGNIFICANT FLIGHT EVENTS, MARINER VENUS 67Event descripti on Time expected,

secLift-off, 0101:00:176 ESTAt la s booster cutoffBooster engine jettisonS t a r t Agena restart timerSus tainer engine cutoffS t a r t Agena primary timerVernier engine cutoffShroud separatio nAtlas-Agena separationAgena engine first ignitionAgena engine thrust at 90 percent chamber pressureAgena engine first cutoffAgena engine second ignitionAgena engine thrust at 90 percent chamber pressureAgena engine second cutoffSpacecraft separationStart yaw maneuverStop yaw maneuverInitiate posigrade rocket

0.0128.0131.0280.1297.3298.7317.4319.5321.5370.7371.9515.21318.11319.31414.91575.11578.11587.11875.11891.0osigrade rocket burnout I .

TABLE IV-11. - ATLAS SUBORBITAL COAST ELLIPSEPARAMETERS, MARINER VENUS 67

ParameterSemimajor axis

Semiminor axis

Radius vector magnitude at apogee

Inert ial velocity at apogee

InclinationPeriod

Unitskmn. mi.kmn. mi.kmn. mi.m/secft/secdegmin

Expected4489.982424.403985.082151.776558.583541.355724.9718 782.729.80249.72

Actual time,sec0.0

128.6131.6281.7296.7308.3317.9320.1322.3380.4381.6525.31319.81320.91415.31576.71579.71588.71876.81892.5

Actual4550.792457.233984.862151.656558.533541.325724.82

18 782.2329.81649.91

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TABLE W-m. - ATLAS INSERTION VELOCITIES AT

unitskmn. mi.kmn. mi.mindeg

VERNIER ENGINE CUTOFF, MARINER VENUS 67

Actuala194.5105.0181.598.088.229.9

ParameterVelocity magnitude

Altitude rate

Lateral velocity

Unitsm/secft/secm/secf t jsecm/secft/sec

Expected5768.9

18 926.8481.981581.3

0.00.0

TABLE IV-IV. - PARKINGORBITPARAMETERS, AGENA

ParameterApogee

Perigee

PeriodInclination

Actual5768.6

18 926.0481.041578.2

1.49 right4.9 right

aSecurity classification regula-tions preclude listing togetheractual and expected parame-ters of Agena.

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TABLE IV-V. - SPACECRAFT TRAJECTORYPARAMETERS AT FINAL INJECTION

ParameterbVis viva energy, c3Radius

Velocity

Flight path angleInclination

units2(km/s ec)

(n. mi./sec) 2kmn. mi.km/secft/sec

Actuala8.61592.512

6569.533547.26

10.9435 892.39

1.9730.31

%ecurity classificati on regulations preclude list-ing together ac tual and expected par ame ters ofAgena.

bDefined using geocentric radius R and iner tialvelocity V of spacecraft as follows:

2 2GMEc - v --where GME is Earth gravitational constant.

R-

18


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Parametera- -B * Ta- -B - R

Radius of clo ses t appr oach

Time elapsed fro m spac ecraft injection to pointof clos est approach, Tf

- - - -

19

U n i t s Actualkm 81 579n. mi. 44 049km -65 311n. mi. -35 265km 75 781n. mi. 4 0 9 1 8hr 3069.932


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TABLE IV-VII. - VENUS MISS DISTANCE PARAMETERS,AGENA (ETR C-BAND DATA)

Parametera- -B - T

Closest approach

Time elapsed from spacecraft injection to pointof clo ses t approach, Tf

Unitskmn. mi.kmn. mi.kmn. mi.hr

Actual211 100113 985

-151 000-81 533231 100124 784

2970.464- - - -aEncounter p arameter s are given in B .T, B e R sys tem which

is defined as follows:-B vector directed fro m center of planet to incomingasymptote of approach hyperbola and perpendicularto it; thus,tance of incoming asymptote

represe nts minimum approach dis-- - - - -B is resolved into components B T and B R where-S unit vector par allel t o incoming asymptote and

referenced to center of planet;perpendicular

and s are-T unit vector lying in ecliptic plane and perpen-

dicular to S-R unit vector completing right-hand orthogonalsystem with s and T (R = s x ?;,

20


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0 4 8 12 16Altitude, km-10 20 30 40 5 0 ~ 1 0 ~Altitude, ftFigure IV-1. -Win ds data at lift-off, MarinerVenus 67.21


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22


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V. ATL AS VEHICLE SYS TEM PERFORMANCEVEHICLE STRUCTURE SYSTEM

by Ri c ha r d T. Ba r r e t tD e s c r i p t i o n

The Atlas st ructu re sys tem cons ists of two major sections: the propellant tanksection and the booster engine section (fig. 111-3). The propellant tank section cons istsof thin-walled, pressure-s tabi lized, stainless-steel monocoque sec tions of welded con-struction. The section is divided by a bulkhead into a fuel (kerosene) tank and an oxi-dizer (liquid oxygen) tank. Skin thicknesses are shown in f igure V - 1 . The maximumallowable differential pre ss ur e between the oxidizer and fuel tanks is limited by thestrength of the interm ediate bulkhead. The fuel tank pr es su re must always be grea te rthan the oxidizer tank pr es su re to prevent re ve rs al of the intermedia te bulkhead. Thetank section is 3 . 0 4 8 meters (10 f t ) in diameter and 1 8 . 5 meters ( 6 0 . 9 f t ) in length.The forward bulkhead is ellipsoidal, and the aft bulkhead is conical. The sus tainerengine is mounted on the thrust cone. Two equipment pods are attached to the s ides ofthe tank.booster engine with two thrust chambers. The booster engine section is attached to athrust ring at the aft end of the fuel tank by a latch mechanism which allows the boosterengine section to be jettisoned.

The booste r engine sectio n consists of protective fairings, a thrust structure, and a

P e r f o r m a n c eThe vehicle str ucture performance was satisfactory. All measu red loads wer e

within the expected limits. The peak longitudinal load factor during Atlas flight was5 . 9 g's at booster engine cutoff. The command to actuate the booster release latchingmechanism was given at T + 131.6 seconds. The mechanism functioned properly, andthe booster engine section jettisoned satisf actoril y.

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Skin thickness/-0.010 (1/2 ard)

/-. 028 1/2/

754.33-783.44-812.55-841.66-870.77-899.88-928.99-

960.0-

Station502522.0542.0562582.0602.06220649.1667.0f

I+(019I.0191.020,-. 24 (3/4 ard)

~ .023

hard)

' t ank

'I1025.0- "-992.5 -1057.5 -

/-Fuel tank

' -- - -IFigure V-1. -Atlas SLV-5 propellant tank section, Mariner Venus 67. (Unle ss notedotherwise, all material is 301 extra-full-hard stainl ess steel. Skin thick ness es andstation num bers are in inches.

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PROPULSlON SYSTEMby Charles H. Kerrigan

DescriptionThe Atlas engine system (fig. V-2) ons ist s of a booster engine, a sustainer engine,

tw o vernie r engines, an engine tank system (pressurization and auxiliary propellant),and an electrical control system. The engines are of the single-burn type. During en-gine start, electrically fir ed pyrotechnic ignite rs are used to ignite the gas generatorpropellants f or driving the turbopumps; and hypergolic ign ite rs are used to ignite thepropellants in the thrust chambers of the booster, sus tain er, and ver nie r engines. Thepropellants are liquid oxygen and RP-1 (kerosene).made up of two gimbaled thrust chambers, prope llant valves, two oxidizer and two fuelturbopumps driven by one gas generator, a lubricating oil syst em, and a heat exchanger.3 3The sustainer engine, rate d 253.5~10 newtons (57x10 lb) thrust at sea level, consistsof a thrust chamber, propellant valves, one oxidizer and one fuel turbopump driven bya gas generator, and a lubricating oil system. The entire sustainer engine system gim-bals. Each vernier engine is rated at 2.98~10 newtons (669 lb) thrust at sea level whensupplied with propellants fro m the sustainer turbopumps during sust ainer engine opera-tion. In the vernier phase of flight, each vernier engine is rated at 2.34~10 newtons(525 lb) thrust at sea level. For this phase, the vernie r engines are supplied with pro-pellants fr om the engine tank sys tem because the sust ainer turbopumps do not operateafter sustainer engine cutoff.

The engine tank system is composed of two sm al l propellant tanks (each approx.51 cm (20 in. ) in diam) and a pressu rizatio n system. This system supplies propellantsfor star tin g the engines and al so fo r ver nie r engine operation after sustainer engine cut-off.

3 3The booster engine, rated at 1468x10 newtons (330x10 lb) thrust at sea level, is

33

PerformanceThe performance of the Atlas propulsion sys tem fo r the Mariner Venus 67 mission

w a s satisfactory . During the engine start phase, valve opening times and star ting se-quence events were within tolerances. The flight performance of the engines w a s evalu-ated by comparing mea sured engine param ete rs with the expected values. These aretabulated in table V-I. All engine cutoff signa ls we re is sued by guidance sys tem com-mands and we re properly executed. Transient s at engine shutdown were normal.

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TABLE V-I. - ATLAS PROPULSION SYSTEM PERFORNIANCE, MARINER VENUS 67

Performance parameter

Duration of booster engine burnDuration of sus tai ner engine burnDuration of ver nie r engine bur n

Performance parameter s

Operating time, se cExpected Actual

128.4 128.6295.5 296.7315.4 317.9

~

Booster engine:Number 1 thrust chamberNumber 2 thrust chamberGas generator chamber pressu re

pressurepressure

Number 1 turbopump speedNumber 2 turbopump speedEngine thrust chamber pres sureEngine gas generator dischargeEngine turbopump speedEngine number 1 hrust chamber

pre ssu re when pump suppliedEngine number 1 thrust chamber

pressure when tank suppliedEngine number 2 thrust chamber

pr ess ur e when pump suppliedEngine number 2 thrust chamber

pressure when tank supplied

%stainer:

pressureVernier:

"Not applicable.

Units

N/cm2ps i 2N/c mps iN/c mps irPmrPmN/c mps iN/cm2psirPmN/cm2psi 2N/cmps iN/cm2ps iN/cm2ps i

2

2

Expectedoperatingrange

386 to 410560 to 595386 to 410560 to 595351 to 382510 to 555

6225 to 64056165 to 6345

469 to 493680 to 715427 to 469620 to 680

10 025 to 10 445172 to 183250 to 265145 to 155210 to 225172 to 183250 to 265145 to 155210 to 225

r + i osec

3925683985 76358519

63166202

483700446648

10 422174252

(a )(a)177

256(a)(a )

Flight values at -3oosterenginecutoff

392568398576358519

62866187

483700441640

10 292173250

(a)(4173

250(a )(a)

ustainerenginecutoff

Ternierenginecutoff

152220(a )(a)152220

26


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I Staging line

T u r b in eexhaust-

Hot gas

Fuelduct-

(a) Booster.F i gu r e V - 2 - At las propuls ion system, Ma r in er Venus 67.

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/ 7-0xidizer bleed valves

Propellant utilization valve. servocontroller-,Head suppression valve \servocontroller-, \

valve

Re ressureressu e

hzTszs Oxidizer, liq uid oxygenFuel, kerosene (RP-1)

Toboosterengines

Vehicle Vehicleoxidizer fueltank tank. .,-Oxidizer regulator

-Turbine exhaust

(b) Su stainer and vernier.Figure V-2. - Concluded.

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PROPELLANT UTI LlZA TION SYSTEMby C l i f f o r d H. A r t h

Desc r i p t i onThe Atlas propellant utilization system (fig. V-3) is designed to cause near simul-

taneous depletion of both propellants. This is a digital type s yste m which adjusts theoperating mixture r ati o of the sus tainer engine by sampling the propellant volume rat ioat six discrete points during flight. Six fuel- and six oxidizer-level sensors are posi-tioned in the propellant tanks so that both se ns or s will uncover simultaneously if thepropellants are being consumed at the proper ratio. Tf the propellant usage r ati o is in-correc t, one sen sor of a pai r will uncover before the other sen sor . The time differencein the uncovering of the sen so rs comp rising a pair is direct ly proportional to the propel-lant usage ratio e rr or. If this time difference is greater than the limit err or times foreach sensor pair, the propellant utilization valve will be commanded to the full open o rclosed position, depending on which sen sor uncovers first. If the actual err or t ime isless than the limit er ro r time, the valve will be commanded to something less than ful lopen or closed position. This adjustment will theoretically res ult in a zero err or t imewhen the liquid leve l reaches the next s ensor pair.

This difference in uncovering time fo r each sen sor pair is measured and an e rr orsignal is transmitted to a hydraulic control unit. This hydraulic control unit directlycontrols the position of the propellant utilization (fuel) valve and ind irectly controls theposit ion of the liquid oxygen valve. When an error signal is sen t to the propellant utili-zation valve fo r an incr eas e in fuel flow, the fuel pump discharge pr ess ur e will decre aseas the valve moves open. The liquid oxygen head suppre ssion servocontro l sen ses thisdecreasing pre ss ur e and causes the liquid oxygeq head suppression valve to move to re-s t r ic t the flow of the liquid oxygen to the thrust chamber , thus decreasing the liquidoxygen injection pr es su re by approximately th e sam e amount as the decrease in RP-(fuel) pump discharge pre ssure. The combined performance of the liquid oxygen headsuppression valve and the propellant utilization syst em r esu lts in a near-constant totalflow weight of propellants to the sustainer engine.

P e r f o r m a n c eThe burnable propellant resid uals in the propellant tanks at sus tainer engine cutoff

wer e 260 kilograms (574 b) of fue l and 207 kilograms (457 lb) of liquid oxygen. Theresiduals would have allowed the susta iner engine to burn a n additional 2 . 3 seconds.

29


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However, the proper velocity had been attained, and the guidance system shut down theengine. If the flight would have continued to theoretical liquid oxygen depletion, the totalfuel remaining would have been 172 kilograms (380 lb).The differential pr es su re between the uthe pre ssu re differential indicates zerosense por t uncovers, a time interval issusta iner engine cutoff. This time interval combinedin determining the propellant residua ls. The fuel sens or por t uncovering time w as2 . 1 seconds, and the liquid oxygen sens or por t uncovering time w a s 6.0 seconds.

Table V-I1 presents the error between the fuel and liquid oxygen sensortimes for all six sensor pairs. The data show that th e err or t imes are al l within thelimit times; thus, at no time during the flight w a s full correc tion capability necessary.

Fuel and oxidizer p res su re sensing ports provide the final

sensor

TABLE V-II. - LEVEL SENSOR ERROR TIMES,Limit

err or t ime,sec

MARINER VENUS 67

--+----1.0188.4

6 4.2

Actuale r ro r time,

se c

1.0.9.3.05.51.0

Firs t sensoruncovered

Liquid oxygenLiquid oxygenLiquid oxygenLiquid oxygenLiquid oxygen

RP-1 (kerosene)

First sensoruncovering

time,

T + 7.2T + 46.2T + 3.8T + 114T + 192.7

30


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E.-Ua

calIE

31


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HYDRAULIC SYSTEMby Eugene J . Cies lew icz

Desc r i ti nTwo hydraulic sy stems, shown in figure V-4, are used to supply fluid power for

operation of the susta ine r contro l valves and fo r thrust vector control of all A t l a s en-gines. One sys tem is used fo r the booster engine and the other fo r the sustainer andvernie r engines.

The booste r hydraulic sy stem provides power solely for gimbaling the two thrustchambers of the booster engine system . System pr es sur e is supplied by a single,pressure-compensated, variable-displacement pump driven by the engine turbopumpaccessory drive . Other components of the sys tem include four servocylinders , a high-pr es sur e relief valve, an accumulator, and a rese rvo ir . Engine gimbaling in responseto flight control commands is accomplished by the servocy linders, which provide sepa-rate pitch, yaw, and ro ll cont rol during the booster phase of flight. The maximumbooster engine gimbal angle capability is * 5O in both the pitch and yaw planes .

The sustainer hydraulic system is si mi la r to that of the booster. It provides hy-draul ic power f or gimbaling the sus tain er engine, for sustainer engine control valvesand fo r gimbaling of the two vernier engines. The sustainer engine is held in the cen-ter ed position until booster engine cutoff. Any disturbances cr eated by the engine differ-ential cutoff impulses are damped by gimbaling the susta ine r and vernie r engines. Thesustainer engine is again centered during booste r engine section jettison. Vehicle ro llcontrol is maintained throughout the sustainer phase by differential gimbaling of the ver-nier engines. During vernie r solo operation, after sustainer engine cutoff, the vernierengines gimbal actuators are provided with hydraulic pr es su re f rom two pressurized ac-cumulators. Actuator limit tr avel of the vernie r engines is *70, and the s ustain er en-gine is *3O.

P e r f o r m a n c eHydraulic system pr essu re data fo r both the booster and sustainer-ve rnier circui ts

are shown in table V-111. Transfer of fluid power fr om ground to airborne sy stems be-fo re lift-off was normal. Starting tran sien ts produced a normal overshoot of about10 percent in the hydraulic pump discharge pre ss ure s. Pre ss ur es, except for the ex-pected transi ents at lift-off, booster engine cutoff, and sustainer engine cutoff, werestable throughout the Atlas flight phase. Gimbaling of the engines was well within thegimbal capabilities and in accordance with the flight control and guidance requirements.32


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TABLE V-III . - ATLAS HYDRAULIC SYSTEM PERFORMANCE DATA,MARINER VENUS 67

Aft-off

21483115217231502124308020823020

Performance parameterBoosterenginecutoff21483115217231502124308020823020

Booster pump discharge pressu re,ab solute

Booster accumulator pressure ,ab solute

Sustainer pump discharge pressureabso ute

Sustainer-vernier pressure,absolute

Units

N/cm2psiN/cm2ps iN/crn2psiN/cm2psi

"Not applicable after booster engine cutoff.bNot applicable after sustainer engine cutoff.

Flight value at -justainerenginecutoff

(a)

( 4

2124308020753010

Iernierenginecutoff

3 3


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Reliefvalve

Riseoff disco nnects

(a) Booster

disconnects (b ) Sustainer.F igure V-4 . -At la s hydraul ic system, Ma r in er Venu s 67.

Rel ie f va lvec 77

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PNEUMATIC SYSTEMby Eugene J. F o u r n e y

D e s c r i p t i o nThe Atlas pneumatic system suppl ies helium gas for tank pressurization and for

various vehicle contro l functions. The system cons ists of th ree independent sub-systems: propellant tank pressurization, engine control, and booster section jettison.The system sch ematic is shown in figure V - 5 .

The propellant tank pressuriz ation subsystem is used t o maintain propellant tankpressures at required levels to support the pr ess ure stabilized tank structure, and tosatisfy the inlet pres su re require ments of the engine turbopumps. In addition, this sub-system supplies helium to p ress uri ze the hydraulic rese rvo ir s and turbopump lubricantstorag e tanks. The subsystem consis ts of six shrouded helium storag e bottles, a heatexchanger, and fuel and oxidizer tank pres sur e regulators and relief valves. The sixshrouded helium storag e bottles with a to ta l capacity of 724 144 cubic cent imeters(44 190 cu in. ) are mounted in the jettisonable booste r engine section. The bottleshrouds are filled with liquid nitrogen during prelaunch operations to chill the helium inord er to provide a maximum storage capacity at an absolute pr es su re of 2068 newtonspe r square cent imete r (3000 psi ). The liquid nitrogen dra ins from the shrouds a t lift-off. During flight, the helium pa ss es through a heat exchanger located in the booster.engine turbine exhaust duct and is heated before being supplied to the tank pres su reregulators. Control of propellant tank pressurizatio n subsystem is switched from theground to the airbor ne re gulators at about T - 60 seconds. The airbor ne regula torsare se t to control fuel tank gage pre ss ure between 4 4 . 1 3 and 46.2 newtons pe r squarecentim eter (64 and 67 psi ) and the oxidizer tank pr es su re between 19.65 and 21.37 new-tons per square centimeter (28.5 and 31.0 psi). Pneumatic regulation of tank pres sur eis terminated at booster staging. Thereafter, the fuel tank pr es sur e decays slowly.The oxidizer tank pres su re is augmented by liquid oxygen boiloff; therefore, the pres-su re decay in this tank is much smaller.

The engine control s subsystem supplies helium pr es su re for actuation of enginecontrol valves, fo r pressurization of the engine start tanks, for purging booster engineturbopump seals, and fo r the refe rence pre ss ure to the regula tors which control oxi-dizer flow to the gas generator. Control pre ssur e in the system is maintained throughAtlas-Centaur separation. These pneumatic require ments are provided from a 76 000-cubic-centimeter (4650-cu-in. ) supply bottle.staging latches fo r jettison of the booster engine section. A command from the Atlas

The booster engine jettison subsyste m supplies pr es sur e to release of the pneumatic

35


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flight cont rol sys tem opens two explosively actuated valves to supply helium press ur e tothe 10 piston-operated staging latches. Helium for the sys tem is suppl ied by a 14 257-cubic-centimeter (870-cu-in. ) bottle charged to a gage pr es su re of 2068 newtons pe rsquare centimeter (3000 psi).

Design range

19.65 to21.3728.5 to 31.044.13 to46.2064.0 to 67.00.35 (min.)0.5 (min.)

2344 (rnax.)3400 (max.)2344 (max.)3400 (max.)----------------------

Performanee

FlT - 1 0 T - 0

seconds secondc

21.83 20.131.8 29.345.71 45.0266.3 65.3

11.72 10.2717.0 14.9

2172 20723150 30052124 20583080 298777.5 79.4-320 -317

The pneumatic system performance was satisfactory. Propellant tank pr es su re swer e satisfactory and all control functions wer e performed properly. Pneumatic sys-tem data are presented in table V-IV. Liquid oxygen tank ullage pr es su re oscillationswere within the range experienced on previous flights. Pr io r to lift-off, oscillation fre-quencies of 3 e 25 hertz were measured. The oscillation amplitudes (differential pr es sur eac ro ss the bulkhead) var ied with a maximum peak-to-peak amplitude of 2 . 0 1 newtons persquare Centimeter ( 3 . 0 psi). After lift-off, these oscillations inc reased in frequency to5 .25 hertz and inc reased in amplitude slightly. These oscillations damped out within ap-proximately the same time span as had been experienced on other simila r flights. Theseoscillations re su lt from the configuration of the regulator and are considered normal.

TABLE V-IV. - PNEUMATIC SYSTEM PERFORMANCE, MARINER VENUS 67Performance parameter

Oxidizer tank ullagepressure, gageFuel tank ullage

pressure, gageIntermediate bulkhead

differential pressure

Sustainer controls botthpressure, absolute

Booster helium bottlepressure, absolute

Booster helium bottletemperature

Ueasure-ment

number

AFlPAF3P

AF116P

AF241P

AF246P

AF247T

units

N/cm2psiN/cm2psiN/cm2psi

N/cm2psiN/cm2psiKO F

it values at -3oosteenginecutoff20.4829.745.365.7

11.9717.5

-17372520434630

43.15-382

justaineienginecutoff

a19. 44a28.2'33.58a48. 713.7920.0

16892450(a)(a )(a)(a)

Vernieienginecutoff

'19.44a28. 2a33. 58a48. 713.7920.0

10621540(a)(a)(a)(a)

Remarks

aHelium supply bottles ar e jettisoned with booster engines at booster engine cutoff +3 seconds; therefore, regulator designrange is no longer a criterion.

36


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Boiloflvalve

w Manua l valveRelief valveRegulatorExplosive valve& Motor valve

I-1 heckvalveOrificed check valve

@ Orifice< ent to atmosphere,-76 000 cm3,-4650 u in.1

Sustainercontrol bottb lubrication tan k-

Closed atT + 20 sec

Booster staging ,-o. 0104 cmdisconnect line I 0.004 in.

Fuel tankHelium pressu rizatioicharge linel ine Liquid nitrogencooled heliu mbottles, 121000cm3 (7365 c u in. )-I

.~lIIIIIlPRiseoff disconne ct panelFigu re V-5. -At las vehicle pneumatic system, Ma rin er Venus 67.

Heliumchargel inen m - m m

Vehicle

8

Ground equipment

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GUIDANCE AND FLIGHT CONTROL SYSTEMby Dean W. B i t l e r an d James L. Swave ly

D e s c r i p t i o nThe Atlas flight path is controlled by two interr elated systems: the flight control

sys tem and the Mod ID[ Radio Guidance System. The flight control sys tem di rec ts thevehicle in a programmed open-loop mode from lift-off through vernier engine cutoff.During the period between T + 100 and T + 110 seconds the Mod I11 Radio GuidanceSystem may generate and transmi t pitch steeri ng signals to the vehicle. During the sus -tainer and vernie r solo phases of flight, the Mod 111 Radio Guidance System may generateand transmit pitch and yaw st eer ing signals to the vehicle. The transm itted steeringsignals a r e received in the airbo rne guidance syst em and routed to the flight controlsystem to provide correc tions for vehicle deviations from the programmed trajectory.

The Mod I11 Radio Guidance System is the primar y source for initiation of d iscr etecommands for booster engine cutoff (BECO), sustainer engine cutoff (SEGO), vernierengine cutoff (VECO), shroud separat ion , start Agena timer (SAT), start Agena restarttim er (SRT), and Atlas-Agena separation.

The Atlas flight control system (fig. V-6) consis ts of the four major componentsdescribed be low:

(1) The displacement gyro canist er contains t hree single-degree-of-freedom,floated, rate integrating (displacement) gyros; one single-degree-of-freedom, floated,rate gyro; and associated electronic c ircuit ry for gain selection and signal amplification.The displacement gyros are mounted in an orthogonal triad configuration alining the in-put axis of each gyro to its respective vehicle ax is of pitch, yaw, or roll. Each displace-ment gyro provides an e lec tr ica l output signal proportional to the difference in angularposition of the mea sured axis from the gyro reference axi s. The input ax is of the rategyro is alined with the vehicle r ol l axis. The ra te gyro provides an e lec tric al outputsignal proportional to the angular rate of rotation of the vehicle about the gyro input(reference) axis.

(2) The rate gyro can ister contains two single-degree-of-freedom, floated, rategyros and associated electronic circuit ry. The input axes of these rate gyros are alinedto their respective vehicle axes of pitch and yaw. Each rate gyro provides an e lectric aloutput signal proportional to the angular rate of rotation of the vehicle about the gyro in-put (reference) axis.tegrate, and algebraical ly sum gyro output and engine position feedback signals. Theelec tri cal outputs of this unit dir ect the hydraulic actuators which, in turn, gimbal theengines to provide thrus t vector control.38

(3) The servoamplifier canis ter contains electronic ci rcuitr y to amplify, filter, in-


45/112

(4 ) The progra mmer canist er contains an electronic timer; arm/s afe switch; high,low, and medium power elect ronic switches; the fixed pitch program; and circuitry tose t the rol l program prio r to launch. The programmer issu es discret e commands toother subsystems.beacon, and decoder; and a ground station comprised of a monopulse X-band (radar)track subsystem, a continuous wave L-band rate subsystem, and a digita l guidance com-puter subsystem. The major functions (fig. V-7) a r e described in the following para-graphs.

The ground tra ck subsystem measures range, azimuth, and elevation and transm itsa composite message- train containing an address code and the coded steering o r discretecommands. When the ad dr ess code of the received signa l is correc t, the airbor ne pulsebeacon tra nsm its a retu rn pulse to the ground station; and the airborne decoder trans-lates the message, and then issues steeri ng signals or discre te commands to the flightcontrol system . The resulting steering outputs from the decoder a r e 400-hertz squarewaves of variable phase and amplitude which a r e transmitted to the autopilot to torquethe appropriate gyro. The gyro torque rate is proportional to the decoder output. Themaximum gyro torque rat e is 2 degrees per second for 100 percent steerin g commands.

The ground ra te subsystem tra nsmi ts two continuous wave signals of different f re -quencies from a single ground antenna. The vehicle-borne rate beacon is interrogatedby the signa ls from the ground subsystem. The ra te beacon transmits a continuous wavesignal at a frequency equal to the arit hme tic average of the frequencies of the receivedsignals. This signal is received by the c ent ral ra te station and two outlying ra te sta-tions. The two-way Doppler shifts and phase relations of the signals as received atthese ground stations a r e used to determine the vehicle range, azimuth, and elevationrates.The position and rat e information fro m the ground trac k and ground rat e subsys temsis sent to the ground computer. The ground computer solves the guidance equationsevery 1/2 second using position and rate information. The ground computer then gener-ate s ste ering and discrete commands which ar e transmitted fro m the computer to theground tra ck subsystem and then to the vehicle.

The ground trac k subsystem conical scan antenna acquires the vehicle during anearly portion of the flight. Once the vehicle is acquired by the conical scan antenna,tracking is automatically switched to the main tr ack antenna which is on the same mountas the conical scan antenna. The ground rat e subsystem antennas a r e electronicallyslaved to the main track antenna.the cube acquisition method, the conical scan antenna is pointed to one of seven pre-determined cubes along the programmed trajectory. If the vehicle is not acqu ired in

The Mod III Radio Guidance System includes the Atlas airborne pulse beacon, rate

The primary method used to acqui re the vehicle is known as cube acquisition. In

39


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the first cube selected, the conical scan antenna may be automatically switched to sub-sequent predetermined cubes. Alternate methods of acquisition are optical tracking orslaving to real-time vehicle coordinates supplied by the Ea ste rn Tes t Range.

Per fo rman ceThe performance of the A t l a s flight contro l was satisfactory . Lift-off tra nsient s in

pitch, yaw, and ro ll were within acceptable limits. The maximum vehicle displacementangles during the lift-off transient s we re 0.9' clockwise in roll , 0.11' up in pitch, and0.28' right in yaw. The ro ll program was initiated at T + 2 seconds and ended at T +1 5 seconds as planned. The progra mmed vehicle ro ll required to achieve the desi redlaunch azimuth was 2.88'. The actual vehicle ro ll from rol l rate gyro data was 3.12',which was within acceptable vehicle rol l requirem ents.times, actual times, and pitch rates fo r each step of the pitch progra m are listed intable V-V. The actual pitch maneuver dispersions we re within acceptable limits.

Maximum dynamic pr es su re occ urred at T + 72.8 seconds after lift-off. Dynamicdisturbances during the period of maximum dynamic pr es su re wer e sm al l and resultedin the booster engines gimbaling a maximum of 1.55' in pitch and 0.48' in yaw. Thisrepresented 31.0 percent of the engine capability in pitch and 9.6 percent in yaw. Thegimbal angles we re within the maximum gimbal angle predictions based on atmosphericdata (wind soundings) taken at T - 0 (see IV . TRAJECTORY AND PERFORMANCE).

Vehicle displacements resul ting fr om booster engine shutdown we re 0.19' down inpitch and 0.51' left in yaw, and were within acceptable limits . The booster enginejettison sequence w a s normal and the result ing sm al l yaw transient was quickly damped.The pitch transient resul ted in an oscillation of 0.166 he rtz with a peak-to-peak ampli-tude of 1. 8 degrees pe r second This oscillation was reduced to within & l o percen t byT + 164.0 seconds. Sustainer steerin g was initiated at T + 138.4 seconds and was nor-mal for this per iod of flight. The vernier solo phase of flight was normal, and steeringcommands during thi s period we re within acceptable limits.from the A t l a s was initiated by rad io guidance. At this time, the vehicle w a s stable inattitude and separat ion was successfully completed.

Postfl ight evaluation of the Mod I11 ground station data indicated that the equipmentperformed sat isfactorily except fo r the problem with cube acquisition by the groundradar. Acquisition in the first cube at approximately T + 60 seconds was not accom-plished because the p reflight Cube 1 elevation setting fo r the ground station track an-tenna was in e rr or .

The pitch program was initiated at T + 15 seconds as planned. The programmed

After completion of the ver nier solo phase of flight, the command to se parate Agena

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A t T + 60 seconds, when acquisition was expected to occur, the received signal atthe ground station w a s abnormally low. When it became evident to the t rack consoleopera tor that acquisition in the first cube would not occur, he switched tooptical tr ack method fo r acquisition. With the aid of the op

conical scan antenna acquired the vehicle at T + 89.7 seconds. The auo monopulse tracking with the main antenna occu

good track data were presented to the computer by T +tion of the vehicle by Mod III ground guidance had no detrimental effect on the mission.A postflight investigation of the acquisition problem revealed that the manual con-

stant setting on the guidance tr ack console fo r Cube 1 elevation w a s set in err or, whichcaused the track antenna to point to a n elevation angle of 88.7' instead of the requ ired48.8' fo r Cube 1. The 88.7' setting was in accordance with the data sheet supplied tothe console operator; however, a tabulation e r r o r had been made during generation ofthe data sheet for the Mariner Venus 67 mission.

Track lock was continuous fro m acquisition in the monopulse mode until T + 410.2seconds, 88.3 seconds after Atlas-Agena separat ion. Track lock was then interm ittentuntil f inal loss of lock occurred at T + 423.6 seconds when the A t l a s w a s at an elevationangle of 1.9' above the horizon. The signal received by the tra ck subsystem from T +98.9 seconds until final lo ss of lock w a s within 3 decibels of the theoretically expectedlevel.data presented to the computer by T + 70.2 seconds. A f t e r a period of 6.0 seconds ofgood rate lock, the signal became intermittent due to the acquisition problem discussedpreviously. The signal level received by the central rate ground station during this6.0-second period averaged -104 dBm (decibel referenced to 1 MW), compared to theexpected -60 o -70 dE3m. Lock at all rate antennas was again accomplished at T + 89.2seconds, and good rate data were presented to the computer by T + 90.7 seconds. Ratelock w a s continuous the rea fter until T + 403.5 seconds. An intermittent period of lockthen occurred followed by final loss of lock at T + 413.5 seconds,and vehicle flight. Following the flight, the guidance program was successfully verifiedbefore remova l of the program from the computer. A simulated re run of the flight indi-cated that no transient er r o r s occurre d during the flight.

Lock at all ra te antennas was accomplished by T + 68.8 seconds, with good rate

The computer subsystem performance w a s sat isfactory throughout the countdown

The pulse beacon automatic gain control (AGC) monitor indicated a received signalof approximately -67 dE3m during intermittent lock from T + 60 seconds until

T + 73.5 seconds, when the signal level reached -48 a m . The airborne received signalstrength reached a maximum s ignal level of -11 dBm at T + 96 seconds, and then gradu-al ly decayed throughout the flight to -32 dBm at Atlas-Agena separation, and to -76.5dBm at T + 409.5 seconds, when lock became intermittent.

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Pulses were missing between T + 101.8 and T + 102.9 seconds and between T +221.8 and T + 124.3 seconds due to momentary loss of signal. This loss is expectedwhen the look angles fr om the vehicle antenna to the ground station coincide with thenulls in the vehicleborne antenna radiation pattern. A momentary loss of signal alsooccur red during the booster staging sequence as a result of signal attenuation caused bythe A t l a s sus tainer exhaust plume.T + 88.5 seconds, as a resu lt of the acquisition problem. The magnetron current moni-to r indicated good beacon response fro m T + 88.5 to T + 410.5 seconds, and intermit-tent beacon response until the magnetron current dropped to zero at T + 426.5 seconds.

The two AGC monitors on the rate beacon indicated that the received signal strengthsvaried about the threshold level from T + 49.5 until T + 87 seconds. From T + 87 sec-onds until approximately T + 391.5 seconds the signal levels were gre ater than -75 dBm.The signa l strengths decreased to the threshold sensitivity of the rece iver, -82 dBm, atapproximately T + 393.5 seconds. The rate beacon phase detecto r and power outputmonitors indicated that the received signals were processed and that the return signalwas transmitted back to the ground station during the period from T + 87 to T + 403seconds.

The steer ing and discrete commands transmitted f rom the ground station we re re-ceived and processed by the vehicleborne decoder. Spurious pitch and yaw commandswere observed prio r to T i-0 seconds, but these commands were inhibited by the air-borne flight control system.

Guidance s tee ring was enabled within the air borne flight control at T + 80 seconds,at which t ime spurious steerin g commands from the decoder were approximately +10 to-8 percent of maximum command (100 percent equals 2 deg/sec gyro torque rate). Thebooster steering resulting fro m these spurious commands was negligible.N o steering commands were transm itted to the vehicle during the period pro-grammed fo r Mod I11 guidance commands between T + 100 and T + 110 seconds becausethe pitch attitude er ro r was less than the computed flight threshold for steering com-mands. The computed flight threshold is a combination of a predetermined pitch attitudeincrement plus an uncertainty increment from the estimation of the "true" attitude. Themagnitude of the uncer tainty increment is dependent upon the amount of tracking datasupplied to the computer. The uncertainty increment at T + 100 seconds for normal ac-quisition, and continuous tracking data, is small. Therefore, the computed flight thresh-old is approximately equal to the predetermined pitch attitude increment, which, fo r thismission, was a l-sigma pitch attitude dispersion.

Because of late acquisition on this flight, the uncertainty increment of the computedflight threshold at T + 100 seconds was larg er than for a normal acquisition. There -fore, the computed flight threshold was greate r. The computed and expected flightthresholds are shown in figure V-8.

The magnetron current monitor indicated intermittent pulse beacon response until

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Also shown in figu re V-8 s the desired booster attitude correction, for the actualtrajectory, that would result in the proper altitude rate at booster engine cutoff. Onlya portion of the desire d booster attitude correction can be made because steeri ng isstopped when the vehicle attitude e r r o r is within the threshold deadband.

Sustainer steeri ng commands s tart ed at T + 138.4 seconds. The largest steeringcommand outputs from the decoder were a yaw left command of 55 percent of maximumsteering at T + 139.5 seconds and a pitch up command of 94 percent at T + 141 seconds(100percent equals 2 deg/sec gyro torque ra te) . Steering commands from the decoderwere reduced to within &10 percent by T + 164.0 seconds, and remained within *lo per-cent until sus ta iner engine cutoff. The amplitude and duration of st eering commands in-dicated normal stee ring by rad io guidance.were a 25 percen t yaw right command which was reduced to within &10 percent in1.5 seconds, and an 80 percent pitch down command that w a s reduced to within &lo per-cent in 3.0 seconds. These commands were within the acceptable limits.

Table V-VI presen ts the times and durations of the discre tes generated by the guid-ance computer.

The la rge st steerin g command outputs from the decoder during vern ier solo phase

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TABLE V-V. - ATLAS PITCH PROGRAM, MARINER VENUS 67

Programmed01.018.848.509.678.806,678.551.382.254.254

Time interval. secActual0.9828.819.4914.6552.819.6552.5733.3276.2457(a)

Programmed0 to 1515 to 3535 to 4545 to 5858 to 7070 to 8282 to 9191 to 105.05 to 120.20 to 128.4128.4 to s ustain er engine cutoff'Cannot be determined.

Actual0 to 1515 to 35.135.1 to 45.045.0 o 58.158.1 to 70.270.2 o 8282 to 9191 to 105.1105.1 o 120120 to 128.4128.4 o sustainer engine cutoff

TABLE V-VI. - MOD 111 RADIO GUIDANCE COMPUTER DISCRETETIMES, MARINER VENUS 67

Flight event

Boos ter engine cutoffStart Agena resta rt timerSustainer engine cutoffStart agena timerVernier engine cutoffShroud separatio nAtlas-Agena separation

Guidance computer generatecdiscrete times,seconds after lift-off

128.427281.704296.607308.374317.854319.927321.927

Discreteduration,

0.497

.570

.497To end

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TProgrammer package

Arm-safeswitch

High powerRoll voltage switches117 riii' i iTimingA

Low powerlogic switchesIlockhousecontrol I. Discretecommands .

lPitChGuidance system

Note;0 Transistor low power switch

Remote ratestation 475gyro package,

torqueramplifier

torquer

Gyro package

displacement signalYaw gyroamplifier

RolltorquerexcitationRol Roll gyro- displacement- ignal -

c gyro amplifierI I '

I -RollIyro packageSMRD logic- itch gyro Pitch Pitch gyrotorquer displacement -t ignal --amplifier ,- gyro amplifier

Pitch rateemote ratestation 415gyro package,

c

FigureVd . -Flig ht control system

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Filter servopackage

integratornetwork

Servo-amplifier

Servo-amplifierodulator

I Filter Iintegratornetwork +Roll

Servo-amplifieremodulator Modulator-Roll

Servo-amplifierI ' I

Hydraulic actuator assemblies

C

;yaw and ro ll

Hydraulic4p controller Rollchannt

Feedback Verniertiransducer ,number 1

, I

ne1

Pitchchannel

2

block diagram, M ari ne r Ven us 67.47


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beacon

Ratesystem

t evar o

.2

U

2L-0LLw-z%0- -.1-

..-Sc.-

-.2

.3 -

-

.l-0-

-

I I1

Flightcontrol -

Track Azimuth -system Elevation -t Commands Computer- -

Wannn ra+n

Figure V-7. - MOD 111 guidance system block diagram.

1-Expected flight thre shol d\\(normal tracking)\ Actualdeadband

I--368 100 102 104 106 108 110 112Time from lift-off, sec

Figur e V-8. - Thres hold for booster steering, Mar ine r Venus 67.

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ELECTRICAL SYSTEMby C l i f f o r d H. A r t h

D e s c r i p t i o nThe A t l a s electrical system supplies and distributes power t ou se r systems. The

elec tri ca l system consists of four 28-volt dc manually activated batteries ; a 115-volt ac ,three-phase, 400-hertz inverter; a power changeover switch; a distribution box; twojunction boxes; and related elect rica l harnesses . The main 28-volt dc batte ry suppliespower to the flight control system, the airborne radio guidance system, the propellantutilization system, the propulsion system, and the inv erte r. Another 28-volt dc batte rysupplies power to the telemetry system, and the remaining two 28-volt dc batter ies sup-ply power to the flight termination sys tem. The inverter supplies power to the flightcontrol, the propellant utilization system, and the airb orne radio guidance system.Phase A of the inverter is used as a phase reference in the flight control and the radioguidance system.The vehicle flight control, propulsion, airborne radio guidance. and propellant uti-lization sys tems operate fr om ground regulated dc and ac power so urces until 2 minutespr io r to lift-off. At this time, the power changeover switch is used to transfer powerfr om ground sources to vehicle electri cal power supply.

P e r f o r m a n c eMeasured electr ica l system parameters were within specifications throughout flight.

Electr ical system performance data a r e shown in table V-VI1 for selected times.

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TABLE V-VII. - ELECTRICAL SYSTEM PERFORMANCE, MARINER VENUS 67Boosterenginecutoff

27.8399.2115.5116.0115.9

Perfo rman ce par amet er Units Specification1ustainelenginecutoff27.9399.5115.5115.9115.8

Main bat ter y voltage, dc I 1 28(I;::Vernierenginecutoff

28.0400.1115.2115.9115.5

Inverter frequency115*0.5115d.7115A.hase C voltage, ac

Atlas-Agenaseparation

28.1400.1115.0115.8115.5

Flight values at -Lift-off

27.9399.2115.5116.2115.9

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TELEMETRY SYSTE Mby Edwin S. J e r i s

D e s c r i p t i o nThe Atlas telemetry system consists of a telemetry package, a manually activated

28-volt dc battery, associated transducer s, wiring harnes s, and two antennas. Fig-ure V-9 shows the teleme try system . Appendix B sum mar izes the launch vehicle in-strumentation.Modulation ( P A M / F M / F M ) telemetry system consists of a transmitter, commutatorassemblies, signa l conditioning components, and su bc ar ri er oscil lators . The telemetrytransmitter has an output power level of 3.5 to 6 w a t t s and requires 28 volts dc for oper-ation. The telemetry sys tem is designed to use standard Interrange InstrumentationGroup (IRIG) subcarrier channels 1 o 18. Channels 2 , 3 , 4 , and 1 2 a r e continuousdirec t, and the transducer prov ides the modulating frequency. Channels 5 to 10 a r econtinuous, using subcar rie r oscil lators . Channel 11 is commutated at 2.5 rps;channel 13 is commutated at 5 rps; channels 15 and 16 are commutated at 10 rps; andchannel 18 is commutated at 30 rps. Channels 1, 14, and 17 were not used for thisflight. The outputs of all subcarrier channels are multiplexed to allow continuous fre-quency modulation of the 249.9-megahertz ca rri er wave.

The 18-channel Pulse Amplitude Modulation/Frequency Modulation/Frequency

P e r f o r m a n c eA t l a s telemetry performance w a s sat isfactory on Mariner Venus 67. One-hundred

and sixteen measurements (table V-VIII) we re made, and all yielded usable data. Notelemetry p roblems occurred during the countdown or during the flight. Signal streng thw a s adequate during flight except fo r the expected 1-second loss of signa l at booster en-gine section jettison. This loss of sign al occurs as a re su lt of signal attenuation by theA t l a s sustainer exhaust plume. Ca rr ier frequency and commutator speeds wer e stable,and no data playback difficulties were encountered.shown in appendix C.

The location of the telem etry stat ions and the telemet ry coverage provided are

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TABLE V-VEL - ATLAS TELEMETRYFlocked-in measurements are either inactive o r will indicate less than 5 percent of

~nalys i sategory Interrange Instrumentation Group1 7 3 4 5 6 7 8 9 10

Continuous

Total number of

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PARAMETERS, MARINER VENUS 67information bandwidth prior to launch. They becolpe active either at engine ignition or liter in flight. ](IRIG)channel number and commutation rate Alternate

15 1610 rps

t616X I3E1D362D363D319ov3384X

34C G280VG82E G282VG3V G287VG279V G2E8V

S252D S259DS253D S26ODS255D S26lDS256D S29OXS257DE29lXS258D

G590V G592VG591V G593V

u112v--p-zg130P H224P4 D7V

I25 19

17

0

measurementsindicative of

30 rps performance18

F1P P116PF3 P

5236x

S53R UlOlA

measurements - 116 153


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Antenna Antennav . . _. _vI pler1 I IT-cou

Figure V-9. -Atl as a ir b r ne telemetry system, Mari ner Venus 67.

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FLIGHT TERM INATION SYS TEMby Edwin S . J e r i s

D e s c r i p t i o nThe A t l a s contains a vehicleborne flight termination system (fig. V-10) which isdesigned to function on receipt of command signals from the ground stations. This sys -

tem includes two receiv ers and two batteries for redundancy, a power control unit, anelectr ical arming unit, and a destructor. The tw o batter ies operate independently ofthe vehicle main power system.

The Atlas flight termination s ystem provides a highly reliable means of shuttingdown the engines only, or shutting down the engines and destroying the vehicle. Whenthe vehicle is destroyed in the event of a flight malfunction, the tank is ruptured with aconical-shaped charge, and the liquid propellants are dispersed. The operation of theflight termination system is under command of the Range Safety Officer only.

P e r f o r m a n c ePerformance of the flight termination system w a s satisfactory . Prelaunch checks

wer e completed without incident. The only measurements telemetered were rece iver 1automatic gain control, engine cutoff, and destruct commands. The automatic gaincontrol measurement on receiver 1 ndicated that the capability to termina te flight w a smaintained throughout powered flight. Minimum signal streng th measured at the re-ceiver on Mariner Venus 67 was 47 microvolts except fo r the expected loss of signa l atbooster engine section jettison. Five microvolts is the minimum required signalstrength for re ceiver operation. Receiver 2 w a s not instrumented. N o flight termina-tion commands were required, nor were any commands inadvertently generated by anyvehicle system.

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-1ADZV automatic Batteries (2) ih

I TelemetryV 1 Destruct 1

Engine Electricalrelay box armin g device Destruct 2_-

Destructor

Manual fuel cutoff

--

I

III

Figure V-10. -At las f l ight terminat ion system, Ma r ine r Venus 67.

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VI. AGENA VEW CLE SYSTEM P E R ~ O R M A N C ECLE STRUCTU

by Robert N. ReinbergerDescription

The Agena vehicle structure system (fig. VI-1) consis ts of four major sections:forward section, tank section, aft section, and booster adapter assembly. Togetherthey provide the aerodynamic shape, st ructural support, and environmental protectionfor the vehicle. The forward section is basically an aluminum str ucture with berylliumand magnesium panels. This section encloses most of the electrical, guidance, andcommunication equipment and provides the mechanical and el ect rical inte rface for thespacecraft adapter and shroud. The tank section consis ts of two integral aluminum pro-pellan t tanks, with a sump below each tank to assure the supply of propellants for enginestarts in space. The aft section consi sts of an engine mounting cone structure and anequipment mounting rack. The magnesium alloy booster adap ter assembly consis ts ofthe bas ic adapter section and the Atlas adapter extension. This assembly supports the

gena and rema ins with the Atlas a fter Atlas-Agena separation.

PerformanceThe measured dynamic environment of the st ructure system was within design limi-

tations. The data are presen ted in appendix D.

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1- StationI 526.00

Retrorocket-,RonctPr

Figure VI -1 . - Agena vehic le s t ructure system, Ma r ine r Venus 67.

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S H R O U D -ADA PTER SYSTEMby C. Robert Finkelstein

DescriptionThe shroud-adapter system, shown in figures VI-2 and VI-3, consis ts of an all-

metal, "over-the-nose' ? shroud and a spacecraft adapter. The shroud provides environ-mental protection fo r the spacec raft prio r t o and during launch. The adapte r is thetransi tion section which a ttaches both the shroud and the spacecr aft to the Agena.

The shroud is 3 98 meters (13.0 f t ) long and weighs 165.8 kilograms (363 lb). Itconsis ts of a 1.66-meter (5.43-ft) d iameter cylindrical section, a 15' half-angle conicalsection, an ogive section, and a 0.72-m eter (2.4-ft) diameter hemispherical nose cap.The nose cap is made of beryllium; all other sections are made of magnesium alloy.The shroud is strengthened by magnesium alloy inte rna l rings to which polished alumi-num liners are attached to protect the spacecraft from the rma l radiation. The shroudis clamped to the spacecraft adapter by a V-band which is tensioned to approximately30 000 newtons (6750 lb). A new, spacecraft ele ctr ica l umbilical door was used on thisshroud. This door is held closed during launch by a spr ing and linkage arrangement.

The spac ecraft adapter is approximately 1.52 meters (5 f t ) in diameter and is 0.21meter (0.70 f t ) high, It is made of aluminum and magnesium and is bolted to the forwardend of the Agena. A magnesium diaphragm attached to the adapter isola tes the shroudcavity fro m the Agena. During ascent , a valve in this diaphragm vents the shroud cavitythrough the Agena. The spacecraf t is clamped to the adapter by a V-band which is ten-sioned to 11 125rt334 newtons (2500rt75 lb).after Atlas vern ie r engine cutoff. At thi s time, Agena elec tri ca l power is used to firetw o pyrotechnically-actuated rel ease devices in the shroud separation V-band. Thefir ing of either release device will effect shroud separation. When the V-band is re-leased, four pai rs of spring loaded pushrods, which are in the shroud and thrust againstthe spacecraft adapter, provide the energy to eject the shroud over the nose of the space-craft at a rela tive velocity of 2 meters pe r second (7 ft/sec).

The temperature in the shroud is controlled; cold air is circulated through an ex-te rn al blanket to maintain an acceptable shroud cavity temperature . This blanket isautomatically removed fro m the shroud at lift-off. The humidity in the shroud is alsocontrolled. A nitrogen purge sys tem introduces dry nitrogen at a flow ra te of approxi-mately 2.26 cubic mete rs pe r hour (80 cu ft/hr) into the shroud cavity near the top, andnitrogen exhausts through the spacecraft adapte r. This purge syst em is manually dis-connected pr io r t o gantry rollback.

Shroud jettison is initiated by a radio guidance di sc ret e approximately 2 seconds

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The shroud-adapter system w a s instrumented with one accelerometer mounted in theadapter, tw o shroud separation switches, two semiconductor strain gages to measuretension in the spacecraft separa tion V-band, a temper ature transducer attached to thediaphragm, and a pre ssu re transducer mounted in the adapter.clean room, and the complete assembly is then transporte d to the launch pad and matedto the Agena. Since the spacecraft ca rr ie s fuel, the clean room w a s als o required to bean explosive-safe area.

The Mariner spacecraft and the shroud are mated to the adap ter in an environmental

During ascent, the shroud compartment pr es su re decayed in the expected manner,indicating proper valve operation.

The temperature measured at the adapter diaphragm remained at a satisfactorylevel throughout the ascent part of the flight. The temperatu re w a s 290 K (62' F) tlift-off, decreased to 283.9 K (51' F) bout 100 seconds af te r lift-off, and increased to287.8 K (58' F) t the tim e of shroud jettison.

Shroud V-band pyrotechnics were fired at T + 320.1 seconds and both separa tionswitches closed, which indicated that shroud jettison was properly initiated. Neitherthe spacecraft axial accelerome ter nor the adapter axial accele romete r showed any un-usual disturbance during shroud jettison, indicating that the shroud did not touch thespacecraft and w a s proper ly ejected. The Agena vehicle w a s stable at this time, and nomeasurable rates in roll, pitch, or yaw developed as a re su lt of shroud jettison.

Spacecraft V-band pyrotechnics were fired at T + 1576.7 seconds. The tensionvalues in both instrumented turnbuckles in the V-band fell t o ze ro tension simultaneously,which indicated that spacecraft separation w a s normal and satisfactory . The velocity ofthe spacecraft relative to the Agena w a s 0.81 me ter pe r second (2.66 ft/sec), which w a sthe predicted value.

During final assembly, the spacecraft V-band w a s tensioned to 11 125rt334 newtons(2500*75 lb). A t this time , the actual tension values measured by the tw o instrumentedturnbuckles in the V-band were 1 392 and 11 236 newtons (2560 and 2525 lb). At space-craft separation, the tension values were 11 040 and 11 984 newtons (248

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Sh r o u dseparationplaneT Ma g n e s iu m a l lo y c o n e\Station 242.25 247.0

_____------_ _ _ _ _ _ _ - -

Bery l l iu m nose caNose cap radiation shield&

Figure VI -2 . - Shroud system, M ar in er Venus 67.

Shroud-,T

----.?IStation \J

i J--- Ii - A g e n a A i

238.60---47- 1-Spacecraft adapter-,support fitting \\( 8 places) \\ \\Spacecraft \ \ ,-Spacecraft-band-, \ \ /

\ \

Figure VI-3. - Spacecraft adapter, M ari ne r Venu s 67.61


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PROPULSION SYSTEMby Robert J. Schroeder

Descr ipt ionThe Agena propulsion system, as shown in figure VI-4, consis ts of a propellant tank

pressurization system, a propellant management syste m, and an engine system. Alsoincluded under the propulsion sy ste m are the Atlas-Agena separa tion system , an Agenaposigrade rocket, and vehicle pyrotechnic devices.

The propellant tank pressurizat ion sys tem consists of a helium supply tank and apyrotechnically operated helium control valve to provide the required propellant tankpressures. Before lift-off, the ullage volume in the propellant tanks is pressurized withhelium fro m a ground supply source . The helium contro l valve is activated aft er Agenaengine first ignition to permit helium gas

atlas-agena performance for the 1967 mariner venus mission 19690020216_1969020216

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