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LTV a€-rF.o ..,Au-rt. e; òtwts.oN P O. BOX øZøt, oÅtLAs,22' TEXAs &v ^ TABLE OF COIYTEI\TTS

Page

INTRODUCTION... 1- 1

MANAGEMENT,... T- 2

MISSIONS 1- 3

THE SCOUT VEHICLE 2- L

THE SCOUT SYSTEMS 3_ 1

THE SCOUT PAYLOAD 4- I

LAUNCH SITES AND GROUND SUPPORT EQUIPMENT. . . 5. 1

THE.SCOUT PERFORMANCE. . . 6- 1

ADVANCED SCoUT PERFoRMANCE 7- L

lÊ Notn, This section identified by orange stripe on edge of pages. II\TTRODIICTIOI\T E

:'lii:ilì -jr'ìr,l LTV Astronautics has been participating in 'it I production and launching of ü:l the development, the Scout solid propellant vehicle for over , t:i four years. During this time, Scout vehicles have been utilized in probe, re-entry and ì ,ti orbital missions including the launching of Explorer IX, the first satellite put in orbit Er¡Ë by a solid propellant booster vehicle.

.J h, This is the second revision of a report which :il was prepared to familiarize users and po- -:x performance r.1., tential users of the Scout and its þ-:l capabilities, It hâs been "up-dated" to reflect :l the improved performance ofthe Scoutvehicle of a new fourth stage motor' il through the use -T the ABL X-258. This report is intended tobe t used for planning purposes only' Complete detailed information relating to the inte- \rìl gration of a specific payload with the Scout .:l vehicle is available upon request.

Information in this report is current as of

rl 1. DECEMBER 1963 i-l

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TheLTVAstronauticsDivisionistheprimevehiclecontractor Fr to NASA for the Scout launch vehicle. In this capacity, LTV Astronautics is responsible for the fabrication of aII interstage structure, including guidance and control systems, instrumen' L tation, heat shield and payload separation devices' Aerospaee' g.orrrrâ équipment, including electronic checkout gear as- w511. a1 F; fabricated sìi itre -ecnãnrcal handling gear and launchers, has been underthisprimevehiclecontractforbothinplantuseandlarrn^cþ site operations. LTV Astronautics is responsibte to NASA,fo! e_ as""*tly, checkout and countdown at the WaIIops Island Station" Included in the services that are provided are payload integration to insure compatibility between the payload and the launch vehicle, EI In general, -q flight planning and vehicle data reduction and analysis' ,.{ ¡' '! Scðut vehicles for the Department of Defense are procured by Command ffir the Space Systems Division of the Air Force Systems Sra. tnrouþtr NRSa. By this method, NASA maintains centralized fr technical direction of configuration, procedures, and checkout which results in over-all efficiency in vehicle production, sche- dules, payload coordination and operation. LTV Astronautics welcomes inquiries on the technical aspects of the Scout vehicle and would be pleased to discuss the potential of Scout launch vèhicles in providing an economical booster for many payload requirements. Inquiries should be directed tol Mr. Fred lM. Randall, Jr. Director - Launch Vehicles

Of -ìiiiÊ:i.;È Mr. Milton-E--- Green Program Manager - Scout

LTV ASTRONAUTICS DIVISION P. O. Box 626? Dallas 22, Texas

t-2 g MISSIOI\TS

The purpose cf Scout vehicles is to boost payloads for orbital, probe and re-entry missions at a lower vehicle cost than is possible by any other means.

This goal has been met through the use of proven and tested concepts, components, and procedures. Design simplicity keyed around the solid propellant motors and a modular concept result in short, simple count- downs and lends flexibility to launch operations.

Complete suþporting systems ernbodying the same principles are available with the vehicles. Launch sites at Wallops Island and the Pacific Missile Range are now available. I:

!x tt;J :l ç TrrE scorrT vEr{rcLE I ,Å The basic Scoutspace researchvehicle is afour- stage guided booster utilizing solid propellant ; :;rl;r,i: rocket motors capable of boosting payloads of ;'rli[ vaxying sizes in orbital, probe and re-entry :1..l missions, The inboard profile, weight break- 4 down, systems, payload capabilities, and per- t formance are detailed on the following pages. PAYLOAD * ß -34.0" t.lrljll SEPARATION SY ST EM ìr,, HEAT SHIELD ALTAIR :ì1¡, SPIN UP MOTORS .n MOTOR -)/ \\\\\\\\\\\\\\\\\\\ SP IN B EARING I \ \ \\\\\\\\\.

GU IDAN C E _il 30" i ANTARES jn MOTOR TRANSITION SECTION "D" '\\\\\\\\\\\\\\ \ '.'... \ \ \ \ .¡ \\ \ \ \ \ HzOz --t \ \\\ \ \ ATTITUDE CONÍROL ,t 3t" I SEPARATION CASTOR DIAPHRAGM i:.' MOTO R !Á APPROX. _ï 72. Fr. .: TRANSITION SECTION "C'' \d tl{l \\\\\\\\\\\:r\ \ \ \ \ \ \ \ \\\\\\\\\\ Hzo. î ATTITUDE CONTROL :- i -1 SEPARATION ,.t UJ 40" DIAPHRAGM rl:.- . . ALGOL TRANSITION SECTION "B'' MOTOR t,t ,i,!J ill J HY DRAUL IC CONTROLS J ET VAN ES \\ \ \\ \\\ \ \\ -\\1\1\r\s-1\\Ì i_'',_, FINS "-i ,tlúJ TIP CONTROLS BASE SECTION "A" r:,Ìì 2-l .t -.¡ wErGr{T TA'BLE çt TY PI CAL

STAGE WEIGHT Payload Fourth Stage - Inert ____19.1J_ Fourth Stage Burnout 79.71 Fourth Stage - Consumed 509.20 Fourth Stage Ignition 588.91 Third Stage - Inert (Includes Spin Motor Grain) 751.7L Third Stage Burnout 1346.62 Third Stage - Consumed 2590.00 Third Stage lgnition 3936.62 Second Stage - Inert (Includes Nose Cone 34" - 25") 22L6.80 Second Stage Burnout -TlffiAz Second Stage - Consumed 7528.00 Second Stage Ignition f36Sr-4--t First Stage - Inert 3234.90 First Stage Burnout 16916.38 '. First Stage - Consumed 2145ri.O0 First Stage Ignition

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'Eí TIIE BASIC SCOI'T \¡ErIICLE

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l. r tr'..f'. ,:,ll ltr . , I l q TI{E SCOIIT MOTORS A'I\TD SYSTEMS a MoroRS. Motor data are presented for each of the scout motors including nominal dimensional and performance characte ristics. a GUIÐANCE AND CONTROL. The systems are designed to have â capability guidance in and control for performing orbitaì, probe and r e - entry missions. . STAGE SEPARATION. Stage separation is positive, clean, and offers no major disturbances. a PowER AND IGNITION. Dualized ignition systems are employed to achieve high retiability. o DESTRUCT SYSTEM. command and automatic destruct capabilities provide positive thrust termination of vehÍcle motors. a INSTRUMENTATION. Performance of vehicle subsystems can be monitored during pre-launch, checkout, countdown and flight,

MOTOR D.A.TA. The summary chart below lists the solid propellant rocket motors utilized in the Scout Launch Vehicle.

sfAGÊ MAKERS MAKERS uNtl NAME NAME NO. NO, NO. 1 AÌgol II A Aerojet Gen. 2 Castor I XM-75 Thiokol XM-3385 27KS- 55,000

3 Ant¿res Allegany _A.B L-X259 - 38DS- i9,000 IIA BatÌ. Lab. .43

4 A-ltair XM-69 .4.1ì egany x-258-Bt Ball- t-ah-

3-1 f i. MOTOR DA'TA çt I

The ALGOL II-A Scout first stage propulsion unit is being produced by the ) Aerojet General Corporation, Sacramento, California. : iI

-f l 33.66 f'- __t i,

ATGOL IIA ',

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't" ft*E Total Impulse - Lb. - ìif Vacuum 5,469,094 tr Éi Specific Impulse - Lb. Sec./Lb. Vacuum 258.29 Burning Time. - Total Sec, 68.2 T .j Thrust - Avg. - Web - Lbs. - S.L. L04,5L7 I L Weight-Total-Lbs. 23,6L0 Weight-FueI-Lbs. rl 2L,174 n. MassRatio-w,/w. Ë',.i p't ilt Nozzle Ëi: Expansion Ratio 7.32 üt! Weight Consumed - Lbs. 2I,457 ff- W ':ìi È- ,l å11:r: Ër5

tÈ EftI,- $ir! % 3-2 'ir¡ ø-..iil g¡..$:'¿l 4; Ll tliJ ALGOL IIA l PERFORMANCE - VACUUM

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l J tr fiia¡ MOTOR DA'TA. çt ft'! I l" I The CASTOR rocket motor, present secoud stage Scout propulsion unit, is manufactured by Thiokot chemical corporation, Redstone Division, Huntsville, 't_[, AIabama.

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,fi- ü't :: cAsïoR r m €.-P;;ri Castor I-E5 tr i"d-_.; : Total Impulse - Lb. - Sec. '+¡ç Vacuum 2,oo2,o7r Specific Impulse - Lb. - Sec. / Lb. Í- ß Vacuum 2'73.2 E"q llÌL Burning Time - Total - Sec. 42.49 d. Thrust - Avg. - Web - Lbs. -Vac. 62,175 L ri$ Weight - Total - Lbs. 8,869 $r: Weight-Fuel-Lbs. 7,328 ho't il.:.Ìl.rl MassRatio-W,/W pt .826 F.:\ r i firiì;; Nozzle Expansion Ratio 15.8 ,rr,rl Weight Consumed - Lbs. 7,434 ffi.'.E ,i.4

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The ANTARES rocket motor (X259), was designed especialiy for the NASA Scout third stage by Allegany Ballistics Laboratory (ABL), a U.Sr.Navry BuOrd facility operated by Hercules Powder Company, Cumberland, Mary-Iandr, The ANTARES X259 is an improved version of the X254.

ANTARES X259

Antares x259-A3

Total Impulse - Lb. - Sec. Vacuum 719,336 Specific Impulse - Lb. - Sec. - Lb Vacuum 281.5 Burning Time - Total - Sec. 36.26 Thrust - Avg. - Web - Lbs. - Vac. 2L,905 Weight-Total-Lbs. 2,795 Weight-Fuei-Lbs. 2,555 Mass Ratio - w-,/WPt .9t4 Nozzle Expansion Ratio 17.93 Weight Consumed - Lbs. 2,580 k'*r #t 'ç l üJ a*r ær e"--fu-@'¡rfuJii'' J"r,*áf tæF *J GJ d nu*[ ffif

o FUEL REMAINING - PouNDS x 1o-o rJÉtsl.lts coOÈ9Èo)co

_a THRUST (VACUUM) - POUNDS X 10 " oræol\''Þo)æorftsFtslJÈ9

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The ALTAIR rocket motor (X-258), Scout System upper-stage propulsion unit, is manufactured by Allegany Ballistics Laboratory (ABL), a U.S. Navy BuOrd facility operated by Hercules Powder Company, Cumberland, Maryland..

ALTAIR X258

Altair x258-81

Total Impulse - Lb. - Sec, Vacuum 14L,729 Specific lmpulse - Lb. - Sec. / Lb. Vacuum 281.9 B urning Time - Total - Sec. 26.0 Thrust - Avg. - Web - Lbs. - Vac. 5,798 Weight-Total-Lbs. 573 Weight-Fuel-Lbs. 503 MassRatio-W,/Wpr .878 1.'r Nozzle Expansion Ratio 25.08 Weight Consumed - Lbs, 509 !:ì fi 3-8 ALTAIR X 258-BI PERFORMANCE -VACUUM

10 L2 L4 16 1B 20 22 24 26 TIME - SECONDS

3-9 F L-

GI..ITDA.NCE A'ND CoNTRíJ^Lq L r tr- t ;f an attitude t The cuidance and control System provides ielfÈ..tli"ç¡åp'tH""-", L-

three orthogonal axes corresponding to pitch' yaw' VrÞ;i¡ti,,,t, ]' a'' probe, re-entry or orbital flight programs' ,t_

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L FIRST STAGE CONTROLS the vehicle is aerody In the lift-off configuration, ,ffi1 of jet vanes i portional control system featuring a combination ,i tip control surfaces operated by hydraulic servo actuators i throughout the entire first stage Uurning,;p..e;1.,0,$ the vehicle tì provide the majority of the control force during. the L,I provide the contiol forcèr iìiì. aerodynamic tip controls all ';Ug.l.r f.r phase following burnout of the first stage. Upper stage control iij,':l qlri follow. ,,iri [{ìi.' ,il,rlt _ri *TË.;.,

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;lE.' *"1B!, i i-î i.,.-, ,h SECOND AND THIRD STAGE CONTROLS Second and third stage control systems are based on the same concept of J operation as the first stage but differ in the method used to generate the control force. The control forces for these two stages are provided by hydrogen t* peroxide reaction jet motors which are operated as an "on-off" system.The s motors are so placed that moments are set up âbout each of the three axes: i pitch, yaw, and roII. The motors are mono-propellant and utilize 90 percent hydrogen peroxide (H2O2). Propellant pressurization is provided by com- ..$ pressed nitrogen (N2) gas. Design characteristics of the thrust motors and '3 H2O2 tankage capacity follow.

40 LB. RËACTION 3 CONTROL MÔTOR H o -¡aNKs (2) 22

N TANKS (2) :s z t.ÐI TOlvÉR 5I DE _g 2 LB. REACTIOÑ I4 LB. ROLL CONTROL MOTOR MOTOR (4) TANKS (ro) tìut ¿ ,Å MOTOR (4) t. SECOND STAGE THIRD STAGE ,'$ SYSTEM ARRANGEMENT SYSTEM ARRANGEMENT FOURTH STAGE SPIN-UP SYSTEM

,bI The fourth stage which includes the payload receives the proper spatial I orientation lrom the control exerted by the first three stages after which it is spin-stabilized by a combination of four impulse spin motcrs. The minia- [ffi turized rocket spin motors are mounted tangentially in the skirt at the base of i¡"- the fourth stage. Spin-up begins approximately 6 seconds prior to 4th stage ',8 ignition. Spin motor arrangement and characteristics follow. Spin rate vs. 4th l stage moment of inertia is graphically shown in payload design parameters i- section. I qr¡¡ sr¡ce HEAT sHtELo

d SFIN MOTO R

ffi TOWER SIDE rB

PlrcH ÀBL zsa Moron -d ÀX ¡S FOURTH STAGE SPIN-UP SYSTEM 3-13 ;--. GIIIDAI\TCE AIYD COI\TTR, OL q

Design CharacterisÈics Second Stage Third Srage

ROCKET MOTORS 500-1b. Moror 40-1b. MotoÌ -ldtlõ rbs a lã1":ìbs Thrust Levels 490 + 30 tbs l4-lb. Motor 40-lb. Morol ^^ ìZ +ÌZìts. i6+6lts g+ t.O t¡s** 44+ 4 lbs 2-lb. Morol 2.2 + (0,8 - ß.4 Totâl Impulse 100% Dury Cycle 25,560 lbs-sec. ¡Ìtel:mittent 2,560 lbs-sec. 21,300 lbs-sec. 2,240 lbs-sec. Spectfic Impuls€ 500* l5s 40# 154 lbf - sec - 40# t65 14# - 179 Itr 2# - t60

Coast Time Probe Probe 5 sec. Re-entry 20 sec'

Re-entry 250 sec. Orbita-l 600 sec. Weight - lbs. 500# - f5.9 tbs max. eâ. 40# - 2,67 lbs, max.ea. 40# - 2.4 lbs max. ea. 14* - 1.96 lbs. max. ea, 2# - 0.80 lbs, max, ea- HYDROCEN PEROXIDE 'IAN¡1{CE CAPACITY Total Impulse F lblsec 100% F;: Dury Cyclc 25.560 2,560 Specific Impulse It¡ -sec I55 |¡1.. lb Ìó0 Weight Hydrogen 178 L7 .8 Pe¡oxide - lhs Volume hylrogen W] ¡r]t l Peroxide Tankage 3840 384 il9-r cu, in. î :::il:l:l I:. l. :i*l or yew and two (2) ¡ou .oro." õ;;;; tr ^^_ì.lllï'^1Tl;-î:." yaw €¡d rwo (2) rorr motors operatii'g Æ rRoll moror l¡tust conrrolil_c-ltg **Roll level during ,n;; .,"g; ;;;;. - -_'""'.;;;;;; or yaw motor thrust cont¡ol d¡rins vehicf.'""""i" { qîf ':l 8,.1 . *"r,lF,ì _l

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tx 'irl l* s The Scout vehicle four solid propellant rocket motors are joined by inter- Stage structures referred to as "transition sections". Eachtransition section d is divided into upper and lower portions at the stage separation plane. iå Frangible "blor¡/-out" diaphragms join the first and second, and the second and .s thirdìtages. The diaphragm forms an internal clampby the threaded periphery 1l that engÀges two structural threaded rings at the separation plane. Blast .õ pr.""r.r." of the upstage motor ruptures the diaphragm, disengaging the [" periphery and allows the stâge to separate. The third and fourth stages are by a ¡omèO ny a "cold-separation" arrangements of springs heldcompressed the flanges, effecting "?.]¡ àIamp retainer flange. Explosive bolt clamps release by sprir-rg loaded ejection force. t_- separation å t* DIAPHRAGM THREADED

! RING ! SECURING .td THREADED 40 EXPLOSIVE A CLAMP I' - PERIPHERY BOLT (4) ...t ,_å \sr EJECTOR SPRING l'o ßz) lmt ;:'ffi :,

1-- rÃltlllr i',ffi *'-'s SPIN r'''.-.ß u'--l \l BEARING

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POWER AND IGNITION SYSTEM LI To achieve high reliability, dualized ignition systems have been employed. Safety features have been designed into the systems to prevent l accidental or premature firing of the rocket motors. Dual squibs are used in all ignitors, dualized. Eachoneofthe dual squibs is in a separate circuit and is connected to a separate battery so that an electrical com- fl t_ ponent failure wiII affect only one circuit. Ignition of the first stage motor is accomplished by a direct electrical signal provided by launch block- F house command. Second, third, and fourth stage ignitions are coptrolled [, by the gtidance program timer. The system provides for the foliowing firing sequence: first stage ignition from blockhouse command, 2nd T stage ignition, 3rd stage ignition and 4th stage heat shield ejection, and 4th stage ignition and spin motors. The same primary power source is utilized by the ignition and destruct systems. tt

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Rlír: e"rl I F;r, Nrl: Þ"i! 3-16 ø-,rl Hlr::. Hil ì çr DES':IR(ICT SYSTEM

The Destruct System provides capabilities for positive thrust termination of the first three stages to avoid a live uncontrolled vehicle impacting on or near the ground launch facilities or in populated areas. The system provides two methods of destruct, the command method operated by RF link and the auto- matÍc method utilizing lanyard and pressure actuated switches. The system is dualized, using two parallel linear charges per motor which can be detonated by either of two separate dual safe-arm units. The system may be armed or safed remotely frorri the blockhouse, Electrical safing is provided for ground operation. Destruct of the vehicle may be accomplished by RF link at the dis- cretion of the range safety officer or by automatic link if the second stage separates prematurely during first stage burning. Destruct capabilities for the payload can be incorporated if desir-ed. A simplified diagram of the system is depicted below.

IST STG.

2ND STG.

LOCK T/M MONITOR

3RD STG,

T,/M MONITOR SHAP ED CHARGES DUAL SAFE/ARM UNITS

SAF E/ARM CONTROL 3 -1? i IIYSTR(IMEIYTATION ç'.

:: equipment n,.' ffi! i:lî' î""ilij:ïrequired to. ;' provide: oo n u îï"#,,1ï .lxî,,;"g*ounã-*åniy:T_1iï::äi;::i::*iï;,, ï:"ï::j:l : :T" T1" " " : " " il;ï,#J""x;.ï:?SJ'î;*:l"xi:":Ëö;,i;å1å:ïiil.få,iïålT,il;ït#J""X;tiil-?ij;'î":"^:t:g"ji;ä":'Jti monitoring or the rlieht,þ.êïr+'.:riì,rl:'ij+ !"":äîi".i"'"#å,:î":.-^::l-::;"iö';'"#:T'1"":'J-ifå'HJTåi#l# jii i"iåî::,:ïîiïì:î",,ï,::;:i",,yli,;'i:¡"ä;Ë":,,äåiiåT."it"T,iäil-+,.".Ëä. ::lü;""å iï/i"# î:ïv i" ã iï i niili i^ii H,iäléäriî ffchannels. ::,JJ;i,g1"*,*j:-"i::l"t,;l;;;"-l##ä;,å,;'ff"Äüiff sienal volt;ô;""ä;ï';ril;"";äi**',ated"iÏijåi PAM,/FM,/rM '!;;1..,.i,i*:oata ", on,/orron,/off reasurements, both tvpes,ã"utypes, are au"i""ïãi;*r,derivËa rìir-an*l,, +-^* r^^ analog an¿'.' i;ïffl"L'"Ï"""""ements,-! mponents such "^0,.,y.¡tit'r,.lr,Ì:iiiiíi system, conirol sv"t.--. enrr rha ¿r^-+-..^¿ ,- , as the guiOancs.i: ririlil i:ü*ii'îlîi".f ,iå"I;:i3lJ,":,rit î::,'r,ll;:*t;ïi,'-"1,1;äå:ï:ïfo. j: ;",;ä J.i: :i il:i,unit *iiïwin be ill,mounted *: :." " ^* Tit' il;: i ::åJJjäff1 iiT,,. ^" i, t¡r" prvi".¿;";iJ";.:::,åi-,

MEASUREMENT Ë: INPUTS {. T ff \.

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\ fr rttt,. I qFr' qTIIE SCOIIT PAYLOAD O PAYLOAD Payload is defined as all instrumentation, mount_ ins strucrures, separatio" il;.;ää,ïTj"]in". atlied hardware attached t" tr,":jå"üiî except li"s" :1to,"", - spin motors, nozzle skirt (when ;o;;;", and ysgd in ir,u"oct"f p.viå"J Jiie"l guidance and control. e aetail ;;;ã;í;ri*:;"tn" standard paytoad comnartment i" p"àriJJ'* in" following pages. Other proven designs *'are avail_avq'r able for special payloao ,"q"i"u_"ii"-. - C VOLUME The payload compartment volume the is dictated by -inside dimenìions or tne rouriñ shield with nominal l',*.n.", clearances between the heat :l::10," motor and payroad. u""t"r uãiu*å.'"îu grven for the payload compartment i*liàu the potentiar paytoad volume ""¿ the. skins of the shields "r"ii"tË¡äiïå., motor tne ãì""ñ'rä;. and the vorume avana¡re"no iir d;;;;äi:, the fourth stage nozzle.

O DESIGN PARAMETERS Design parameters and environment factors for . the paytoad are detaited .;;;;;r;ï;;*;i , including payload separation,"" Oynamic siabili_ ì illl":,.payload thermal enviro,.,ment,-rná *"_ cnanrcal environments of acceleration bration. and vi_

O PAYLOAÐ INTEGRATION LTV Astronautics assists in the integration of payload requirements within the partments "pp"åp1iJ" including essential mechanical"o*_ and electrical interface arran gements.

TYPICAL PAYLOAD

4-r ll

PAYLOAD CONFTGTTRATTON q MoroR,) -(x258 i STANDARD 34,, HEAT SHIELD

The pertinent details ofthis payload compartment are shown on the adjacent page. When Payload and 4th stage separation is not re- : quired, additional payload volume is available. The heat shield !!r which dictates the compartment is of fiberglas honeycomb sandwich construction with a stainless steel nose cap. It is designed to carry fourth stage vehicle Ioads and to keep compartment temperatures within specified limits. Abumper between the heat shield and motor - keeps the payload from hitting the heat shield under high load con- i ditions and transfers lateral inertial loads fromthe motor case to - the heat shield.

The heat shield is fabricated in two half shells with a stainless a ¡ steel nose cap attached to the tower side of the shell and with Í steel half-rings at the base to accept ajoining clamp. Shell restraint iÌ is provided through a series of over center 'latches along the

separation plane and a joining clamp at the base. The latches {. and clamp are triggered and released by drawbars attached to ¡Ì bellcranks and a ballistic actuator near the forward end of the heat shield. Upon release of the latches and clamp, Iatch contained springs force the heat shield apart. The separation plane lies along the pitch axis.

I trI ; ¡f iì rt

hl T

!, i' rit :

ñ ',

4-2 \t:- .r(b T I t__ r\___-.\-r --\.- ;in:-:n 6À rL -i Ì:i, ..i \ É,r*!..á L ,j, I rïã PAYLOAD EXTENDED INTO THIS L.-9 AREA REOUIRES DETAIL COOR_ ffi DINATION WITH SEPARAfION INSTI ,lr I rrq L tj r ..g Ì rì{ rirI

F LÌ@

ì L \$l HEAT SHIELD WEIGHT I r3 233 LBS i :r3

NOSE CONE ,ffß i råJ __t HAROWARE PROTRUDE INSIDE CONTOUR I.I2 INCHES. DOOR LOCATION REOUIRES COORDINI 'il"$ WITH SPECIFIC PAYLOADS. i

.itJ .

-\.-

>--

RANGE SIOÉ

,9AO EXTENDEO 'REOUIRES INTO THIS i, DETAIL COOR_ \. I.'ION WITH SEPARATION INSTI .óo ÍHICK ----\ HEAT SHIELO

25.70 0tA

TOWER SIDE \

4TH STAGE ¿..-'..'.. SEPARA'TI ON PLANE --/ -/' -- ,/ PAYLOAD SPACE AVATLABLE REOUIRES DETAIL COORDINATION WITH LANGLEY RESEARCH CENTER

NOTE- UMBILICAL DOORS ANO RELATEO HARDWARE PROTRUDE INSIOE CONTOUR I .I2 INCHES. DOOR LOCATION REOUIRES COORDINATION WITH SPECIFIC PAYLOADS. 4-3 PAYLO.A.D DESIGN PARA.METERS E

PAYLOAD MOUNTING & SEPARATION

A typical payload mounting and sepa- debris consistent with the configuration system is shown here. The ration requirements of the payloacl. system provides the f ollowing features: Seppration occurs with ignition of the explosive bolts which releases O Minimum tipping of the payload the retaining clamp and allows ejection of the payload and attached O Adaptability to different payload mounting ring. The compression configurations and we ights plate and conical ejection spring are cable-connected to the payload O High reliability - functionally support structure, proven Connectors can be provided across the separation plane to provide elec - O Careful control of the two trical functions between payload and clamp-retained flange surfaces the lower stages of the vehicle, by utilization of payload adapter mounting ring A new, Iighter weight, more sophis- The separation system is designed ticated separation system is under to imparta 6ft./sec. total separation development and will be made avail- veiocity to a 1?5 pound payload with able as soon as it has been flight containment of separation system proven.

! I I t PAYLOAÐ SEPARATION SYSTEM I \.;

4-4 t1 ,Ê EXPLOSIVE BOLT

RETAINING CLAMP

PIN -

- YAW TOWEI

PITCH AXIS VIEW TOOKING AFT ,s:,

-,- SPECIAL BoLr "^ -w- þ',' *\-__-).r' \\ --7 SECTION B-B lt I

-a. I

SECTION .4,-A

PAYLOAD P.å.YLOAD SEPARATTON SySTEryl

vEHrqlg_qENIE& I,ILE

CONICAL EJECTION SPRING

4-5 -nBI PA'YLO.A.D DESIGN PAR,A.METER,S ç' ,.tiM, H DYNAMIC STABILIZATION :{ tìiltt:gÉ The final stage rocket and payload are stabilized for directional stability by ,ffi spin balancing, both statically and dynamically, to minimize flight path errors .tg]' during last stage burning. IT¡' idT Dynamic balancing is accomplished during vehicle stage assembly in the spin , r&Si Fl bâlance facility at the launch site. Spin is clockwise as viewed from booster i{, base to nose. ffi The stage is assembled on the dynamic balancing machine starting with the r transition section. Static and dynamic balancing is performed on each ,',L sequentially attached part. The assembly spin rateis 160+20 H rpm. Temporary p'. ballast used the assembly balance process is in and finally consolidated into t:'

two masses located at a convenient plane on the payload and on a plane near LF the spin bearing on the support skirt. Hr ';*,:l If despin is required for a mission, it can be accomplished by the user or by +, LTV Astronautics, as desired. tr1 f: ",*

#x Ê;i" h tã $li!

L. ¡1! !|j S ra''vr-.oaD DESIGI\T PA.n,.A.METER,S

SPIN ENVIRONMENT The spin environment for various sPIN RATE VS 4TH STAGE MOMENT OF INERTIA VÀCUUM fourth stage total moments of inertia is shown in the adjacent graph for vehicles incorporating the cold (spring e jection) separation system. For total roll moment of inertia, add payload roll inertia to results inthe Roll Inertia Table below. The effect of spin rate on final stage ftightpath deviation indicates a desired minimum of 140 RpM. The spin rates shown in the adjacent graph are those which wiII occur at fourth stage ignition. As a result of the internal gas dynamics, the spin rate at fourth stage burnout wiII be approximately lLls greater than those at ignition. The table shows typical Scout roll moments of inertia for a fourth stage.

W INERTIA OF SCOUT

Propellant - (Consumed) 509.2 2L,3L2 Inert Motor - B.O. 63.8 2,78O Skirt "D" L 5.91 1,56? Spin Motor (Inert) 2.16 2t6 Spin Table 14.85 1, ?93 Grain 1.0 100 Wiring at 368 Separation Clamp 6.20 878 Reflective Tape .2L L7 Payload Separation 15.40 602 629.93 29,633 = 6.4 stug/tt.z

4-7 PA'YLOA'D DESIGN PARAMETERS ç

UMBILICAL PROVISIONS

Although payload umbilical doors may be positioned within a horizontal angular location +15'of the yaw axis centerline on the tower side, PaY- loads are requested to use toP centerline Yaw axis Positive for a more reliable umbilical operation. Vertical Positioning is normallY within the area tretween sta. 51.0 and sta, 22.0 as shown in the adjacent illustration. The doors ånd related hardware protrude inside the contour 1,12 inches. Standard door dimensions are 5 x 5 inches.

The umbilical electrical connectors feature Bendix "twist/pull" quick disconnect plugs with modifications to provide reliable lanyard uncouple. Shock pull-off is set at 45 +5 pounds'

Heating or cooling facilities can be incorporated for temperature control of the paYload, usable uP to launch. Entrance of conditioning air supplied from the launch comPlex unit enters at the same angular Iocation as the umtrilical, with disconnect either prior to or simultane - ous with the umbilical connector re - traction. The cooling sYstem will provide humidity controlled air at selected temPerature between am- bient and 40'F.

4-B åx) P^A'YLOA'D EN\¡IR,OI\TMENT THERIll\AL

The heat shield and stainless steel nose cap protects the payload from aerodynamic heating during ascent and carries fourth stage vehicle loads for all launch angles from a minimum of 78" to nearly vertical. Added protection is afforded by a gold coated radiation shield located on a station-plane at the aft end of the nose cap.

Nose cap and heat shield temperature vs. time are shown on the following page. Heat shield temperature gradient vs. time forboth cylindrical and ãonicat sections are depicted on page 4-11 and 4-12. The trajectory below is used for each of the Plots.

SCOUT ALTITUDE AND MACH NUMBER VS TIME oâo

:lî

Ii TIME - SECONDS 4-9 PA'YLOA'D ENT\/IRONMENT E NOSE CAP AND HEAI SHIEID TEMPERATURE VS. TI'V\E

NosE caP\ fî:1"'"".".

STAGNATION *A' >tr

\ 0 ff re.vr.oe.D ENvrn,oI\TMENT HEAT SHIELD TE'YlPERATURE GRADIENT VS. TI/IAE

CYL IN DRICAL

OUTSIDE - .05 FIBERGLAS

ilililrilililIililIililt I .50 HONEYCOMB CORE ll lilililtl ililililililr I r00 120 rNsrDE - .oó FTBERGLAS *1 150 -l LL H EAT SHI ELD CONSTRUCTION A I z t.Lt ¡¡O É. JV (J f ttl F u ¡ tu L! (L = ul F = 150 F o 120 oz u 100 ul

I LU Ê

.05:..+ .50 .05 -FIBERGLAS HONEYCOMB FJBERGLAS

OUTSIDE SURFACE INSID E SURFACE

SCOUT HEAT SHIELD TEMPERATURE GRADIENTS CYLINDRICAL SECTION LA = 83"

4- 11 PAYLO.A'D ENT\¡IR()NMEI\TT Y HEAT SHIELD TE'YlPERATURE GRADIENI VS. TIIYIE

CONICAL SECTION

__'--] OUTSIDE - .05 FIBERGLAS I t30 IL t50 I I00 COR E ul 120 o t z f o l--- LU É. INSIDE _ .Oó FIBERGLAS LIJ fL I tJ-t =UJ H EAT SHI ELD CONSTRUCTION F =F

IL o I z uJ o ú. U f L! F Í I ut t].i o- ¿. t-- =L!¡ t- 70 ó0 20

.05 ------_ .05 .50 FIBERGLAS FIBERGLAS HONEYCOMB

OUTSIDE SURFACE INSIDE SLJR FACE ---_

SCOUT HEAT SHIELD TEMPERATURE GRADIENTS CONICAL SECTION LA : 83"

4.t2 çr P.A'YL o aD Er\TvrRoNM Er\TT IvIECHANIcAL This section provides a summary of minimum recommended mechanical test requirements for the environmental testing ofbothprototype (qualification test and development units) and fLight spacecraft to be boosted by Scout vehicles. It covers only the mechanical environment exposures prior to pay- Ioad separation and/or fourth stage motor burnout and applies to the complete spacecraft assembly. These basic requirements are considered to be a minimum for the demonstration of the structural, mechanical, and electronic integrity of the spacecraft during flight ascent. I i DYNAMIC BALANCE

The spacecraft in its fourth stage spin-up configuration and with systems r non-operating shall be statically and dynamically balanced, prior to the other I s tests, within the following limits: H

Maximum static unbalance: L2 oz. ín. fi

ü Maximum dynamic unbalance: 200 oz. in. Ë The prototype shall be dynamically balanced at L-I/Z times the maximum ä flight spin rate" Flight hardware shall be dynamically balanced at the maxi- A mum flight spin rate. H ä ACCELERATION ill

H The prototype spacecraft shall be operating -il during the acceleration tests. ïi H in each of three axes. Acceleration applied along the thrust axis should be ffi m I L .L/2 times the maximum calculated level for the spacecraft weight and center llJ ffi . of gravity (CG) for a duration of 3-minutes. Acceleration applied in each of ffi t- F] ffi .,-ì the transverse axes should be at 39 for a duration of l-minute. Acceleration il gradient from the CG should not be overplus or minus 1070. Acceleration test ffi of the flight spacecraft is performed only if deemed otherv¿ise necessary. ffi SHOCK ffi The spacecraft in fourth stage ignition configuration shall be operating ffi ffi during exposure to the shock tests in the thrust axis. The prototype space- w ffiì craft shall be subjected to three half sine pulses of 30g peak amplitude and üg #ffi 10-15 milliseconds totâI duration. The flight spacecraft shall be subjected ffi ffi to a haif sine pulse of 20g peak amplitude and 10-15 millisecond total ffi duration. 4-13 ffi PA'YLOA'D ENVIRONMENT ç' MECHANICAt

VIBRATION: The spacecraft (this includes the support and separation assembly structure which attaches the payload to the fourth-stage. motor) should be operating during exposure to the vibration test. The test levels apply at the interface of the forward motor shoulder of the fourth stage motor. The tests should be performed with a structurally and dynamically similar separation assembly installed. The sinusoidal vibration test shall be conducted by sweeping at a logarithmic rate from the lowest to the highest frequency once for each range specified. The random vibration test shall be of random peak accelerations over the frequency ranges specified. The peak accelerations which exceed the g rms value by more than three times, shall be clipped and shall not be imposed on the spacecraft. Prototype Spacecraft - (This simulates L50Voof. the expectedflight vibration tevel.) Ç' re,xr-. o aD ENvrn,o trIM ENT ,ìAECHANICAL Flight Spacecraft (Flight Acceptance Tests) - This simulates 10070 the expected flight vibration level.

Note 1: The acceleration at the spacecraÏt center of gravity (CG) shall be limited to r3g from one-half (L/2) tlrre first resonant frequency of the spacecraft to one and one-half (l-L/z) the frequency when vibrated in either of the transverse axes.

4-L5 tfii 1r rlÍï, li 'rilffiÏffur scorrr LArrNTcr{ srrEs lr- ....,..',... i AI\D GR,OTIND SIIPPORT l: :i EQTIIPMEl\TT I il .,,' iì Sior¡t launch sites have been established at Wallops Island, Virginia"and Pacific Missile Range (PMR). Experienced launch operations personnel .are available and are currently conducting or assisting in Scout launch operations.

Launch and Ground Support Equipment including the launch tower, receiving inspection and test equipment, assembly and systems test equipment, checkout and launch equipment, and special test equipment are supplied by LTV Astronautics.

PMR LAUNCH SITE POINT A RGU E LLO, CÀLIF.

WALLOPS LAUNCH SITE WALLOPS ISLAND, VIRGIN IÄ LTV ASTRONAUTICS. DALLAS, TEXAS

5-1 PA.CTFTC MTSSTLE R,.A.NGE çr TAUNCH COMPLEX

POINT ARGUELIO, CALIF.

i r l

t- ü fr

F a I

tl e lr!!

¡r-$

aiF-, ß,

ffi,.

I i r.Í-iä i Ç wa.r-.r,oPs rSL.A'I\TD ,i{ TAUNCH COtvtprEx '4 .å +i _r l I WATLOPS ISIAND, -":r VA. fl

._ii-{ I f

,ä ._g I +{j

3iã -è r "\--:l .\Þç' .j gÞ- !I(E|AT.TI MORE i ÞlP i çANNAPOLtS, t- wasu¡Ncro¡.,¡,1 ã """.*;' , I tÅ ì{-\ I ^i'""'^\, ,,{ .T re il

tl Ø' .'- {fl '¡iå '/ .t-" -r ry, Þ iå

,.'d ;Ë r

s,8 *- J ¡ù. 5-3 du f'i,tg TOWER LAtTNCTTER çt WALTOPS ISLAND, VA.

REMARKS: VEHICLE CHECKOUT IN A VERTICALPOSITION. SERYICE PLATFORMS CONTAIN ENVIRONMENTAL PROTECTION FOR VEHICLE.

TYPE: LAUNCH BEAM WITH FITTINGS FOR FORWARD AND AFT ENDS OF f ALGOL MOTOR STAGE.

ADJUSTMENT: f^ AZIMUTH: FROM 90. TO 1BO" TRUE HEADING. ELEVATION: FROM ?0. TO 90" FROM HORIZONTAL. Í

I ;

0-: í!.

ql 5-4 rq! Ç crrncKotrr c o¡[s oLE-BLo crrrror;s E WAILOPS ISLAND, VA.

5-5 tloF"rzo NTI\L CrrE c'Kc)(TT L Arrr\T crrER ç'

POINT ARGUELLO, CAtIF.

WATLOPS ISLAND, VA. (MK II LAUNCHER

REMARKS: VETIICLE ASSEMBLED AND CHECKED OUT HORIZONTAL BEAM ELECTROMECHANICALLY ERECTED.

TYPE: LAUNCH BEAM (ZERO LENGTH) WITH FITTINGS ON FORE AND AFT END OF ALC.OL MOTOR STAGE.

ADJUSTMENT: AZIMUTH: 160" TO 3OO. - PMR 65. TO 205" - WALLOPS ISLAND ELEVATION: ?0" TO 90" FROM HORIZONTAL'

RE MOVAËlLE SHELfER

fRAN 5P ORTE R :;'rl-li'li | ' :' 'iiJ!:rr '.. ìl Hó'Rrz oñt ¿.r-. c rr-'F ör our L ¿'IIN c rr-ER

POINT ARGUEL[O' CAIIF.

WATLOPS ISIAND, VA' (MK II LAUNCHER

RE MOV ABLE SHELfÊR

VERTICAL (LAUNCH) POSITION hs gTITE SCOIIT :ä PER,FOR,MAIYCE

or the sc out as a rroos iä *:t""ætårå?iotttties te r ror o rbital, p j:1"*;;.;;:""ilìi",'_t,i-""ì1",r*"ìir,::î,ï:lilii,i:";;:*,1,*;üg" ro be, lzero angle- of -attack),i":"""ao"n ð ä;ñ"ffh" payroad t"o"' *p'¡iritiã" to the desi¡ea in_ i::i:ffi,"*tåtîîî "ï;ä;"" based onnominar sorid- OR,BITAL MISSIOI\I The performance of the,scout as an orbitar booster, shown ing pages, iltustrates tte..pavtoaá c;;;iîiïu. on the folrow- circurar and elliptical of the vehicle for both pMR" 0rbits when ra*ãirã,i""*uothvr'altops The effects of orbital i""rir"iioïàîiåïro"a, Island and "#. a typicar sequence :ï.;:::l;,Ïå iT,',:i: "' "r''" "äi"îiîiiö sta ge i mp ac t poi,itï "na

3RD STAGE

2NÞ STAGE

lri: I

..i:.r:Ì.1 lST STAGE

],Tì.ìr: CIRCfTLA'R ORBITA'L PERF'ORMAI\T CE q

lr00

1000

900

800

700

ó00 ELTIPTIC.â.L OF,BTTAL PERFORM.A'NCE

POTAR ORBIT

10000 9000 8000 7000 ó000 5000

4000

+z 2000 I ôt¡J l F F 1000 J 900 800 uJ 700 l¡J (9 ó00 o TL 500

400

6-3 ELLIPTICA,.L ORBIT.A,.L PERFOR,MAI\T CE q

EASTERLY ORBIT

r 0000 9000 8000 7000 ó000 5000

4000

3000

2000

I000 900 800

6-4 € EF FECT-L.A'UI\TCFIIhIG OTIIER, TIIAN DTIE EA.ST

CIRCULAR ORBIT CAPABILITY PMR LAUNCH 30

CJ UJ Õ 50 I z o 60 F z J (Jz J F a oe.

PAYLOAD - L BS,

CI RCULAR ORBIT CAPABILITY WALLOPS ISL AN D LAUNCH

.80 o ôTU I =g uo F z -J zo40 J F- õ .,n tt" o

PAYLOAD - L BS.

6-5 Xri .ç,

Fl t. TYPTCAL ORBTTAL-ASCENT çt f;,'ì

¡1. TRA.JECTORY ( þ r !- t. I ): r ii r I

¡¡¡ i l LAUNcH Àzrmutn - 90 DEG' PAYLOAD - 150 LBS. inuucg sITE - wALLoPS ISLAND ì HEAT SHIELD DIAMETER - 34 IN. iÁuNcn ELEVATIoN ANGLE - 84DEG' 1 i. TYPICAL SEQI'ENCE OF E\¡ENTS i

{

6-6 çr TYPICA'L BoosT PER,FOR,MAT\TCE I}IIRING ORBITA'L A'SCENTT

PAYLOAD - IóO LB. EAST LÀUNCH FROM WALLOPS ISLAND

¡, t2 oð t0 oÉ, t- I o tt 6 â 4 o J 2 z 0 æ0 400 500 TIME - SECONDS

NOTE: VEHICLE - (ALGOL llA, CASTOR, X-259, X-258)

.25 u lu

F F u- n20 o o o o I o I o) Èrs I ú t¡¡ o ô J 3 u.¡ F tr -_, lo J

300 ,100 500 TIME - SECONDS

6-7 TYPTC.A'L GROTIND TR.A'CIK, q IMPACT POIIYTS .A.I\TD DISPERSIONS

GROUND TRACK AND IMPACT POINTS

-l EFFEcrrvE LAUNcH ANGLE = 80. B8o "o"'"1'= '' i tou*"" -o-R-T- E

' IST SIAGE _r_c- 2"" slneg eoos.reR rvplcr T- L-A-]r-+T \ I ¡ s"'srÀee eoosrER ¡MpAcr

( sEE DISPERSToN c

l0Ë {q

¡- TYPICAL IMPACT DISPERSIONS I Booster impact dispersion areas are elliptical in shape. Typicai 3a dispersions for orbital boost trajectories are shown in the table. I

DISPERSION AREA DIMENSIONS {-t Booster Length width î_ Stage N. Mi. N. Mi. r I 1 t2 5 ï 2 54 20 É: J 160 68 Í

'(-f çr FROBE MrSSrOr\T

Capabilities of Scout as a probe booster are defined ín the following pages, AII four stages of the vehicle are fired in sequence and the payload coasts to apogee. Peak altitudes, time in a weightless condition and trajectories are provided for the fuII span of launch angles. Typical trajectory characteristics are shown. Complexity of payload re-entry conditions for probe missions depends on peak altitude and velocity, re- entry angle and resulting heating and deceleration conditions.

2 / / ^â rGNtÏoN / O BU RNOU T / / / I/

4TH STAGE

+-3RD STAGE NNNNN -2ND STAGE\\ N\\\NN

6-9 PROBE PERF'OR,MANCE çt DEEP SPACE I PAYTOAD VS. ALTITUDE f Fì I I

I¡

!". I t

NOT E5: I. LAUNCH FROM WALLOPS IS. F' I 2. LAUNCH AZIMUTH IIO DEG. I 3. HEAT SHIELD DIAMETER 34 I F" 10000 f 9000 8000

7000 I i. ó000 5000 t- I z< 1000

I sooo Hl F F

r000 tñ I 900 F..t 800 700 ó00 Ç 500 Í i 400 :i,

300 'Ii .l

ft 200 PAYLOAD L 85. - '

If È!l'!

r- 6-10 t

ì çr PROBE PERFORMAI\TCE TIME AT ZERO t.G,,

6-11 PROBE PER,FORMANCE çt RANGE VS APOGEE

WALLOPS ISLAND LAUNCH - l IO DEGREES AZIMUTH

1000

500 3000 4000

6-t2 4 pnosE PERFoRMArrTcE BOOST TFIROIIGII FOIIRTII STA'GE

LAUNCH AZIMUTH - 90 DEG. PAYLOAD - I5O LBS. LAUNCH FROM WALLOPS ISLAND LAUNCH ELEVATION ANGLE - 84 DEG. HEAT SHIELD DIAMETER _ 34 IN.

ÄLTITUDE. VELOCITY, & LOÄD FACTOR VS TIME

? z 9 t-

r00 r50 200 TIME - SECONDS vEHrcLE - (ALGOL Ä, CÄSTOR, X259, X258)

_-å

j "¡ _L ¡ il ;"d 6-13 _tI $ fll ¡T

RE-ENTRY MrssroN q t)

¡-1

Re-entry performance of the scout follov/ing a gravity-turn are provided in the peak following p"g"s. The booster fires the first two stages, coasts beyond !5' payload back into the altitude and then fires the upper two stages to drive the ì atmosphere. Re-entry capabilities shown should be considered typical of scout capabilities since significant changes in re-entry conditions can be obtained i- { thiough selection of pitch rates, staging times and final stage attitude at ]' ignitión. Increased re-entry velocity may be developed through the use of a re-entrY. f fifth stage during I

I' i

tï- k &

lrF r !. 1".1

i : .tl

YIÀLLOPS ISLAND

6-t4

t:i ¡ Ç nn-TNTRY PERFoRMa.T\TcE A'T FOTIRTFI STAGE BTIRNOTJT PAYTOAD VS. ATTITUDE

25 ',26 27 28 29 RELATìvE vELoctry - rt./sgc. x lo-3

6-15 ú!

rfù E. PERF'ORMANCE AT q rí, R,E-EI\TTRY q, FOTIRTIT STAGE BIIRNOIIT r.r. r PAYLOAD VS. VETOCITY L--

t I

t!" i:--

IF' t.

ll fi l[r, ïi'¡ {

ra. :e-

'"|

10 ,)o

ì E:- çr RE-ENTIÌY PERFORM.A'NCE ATTITUDE VS. RANGE

LAUNCH FROM WALLOPS ISLAND _ RE-ENTRY - RELATIVE VELOCITY 22OOO TO 33OOO FPS RE_ENTRY ANGLE _5 TO _15 DEGREES

600 iii;iiiì.:$i,,r ji¡ÌjtjlÌi;lrj irl ij;.rrt :l;iriai ì' ,::t,tì,:t:::. riirtr',1:: ! *:,i!ì'.li,\i.:t..,. É,Ì;":lr;iili.!t,irt,,.i 500

RANGE - N. MI

6 -17 € ppnron,MAr\Tcr ADVANCED SCOUT

Significant orbital, probe, and re-entry performance im_ provements are achieved with theuseof improved motors which are under development or being planned for de_ velopment,

This program currently has under development an im_ proved second stage motor which is a *oãifi"d version of the Castor motor presently being used. Improvements to the other motor stages have alreadybeen inãorporated. The ABL X-258 recenfly replaced the X_24g as the standard fourth stage motor. Other changes have been the incorporation of a new first stage motor, the Algol IIA, which was a development of the Algol ID originally used, and the X-259 which was a development of the X_Zb4 originally used as the Scout third stage motor.

Prior to detail planning to take advantage ofthe increased performance and payload capability resulting from the use of improved motors, inquiries should be made to determine the time phasing of these developments into the standard Scout vehicle

7-I € epv¿,.¡[cED scorrr CIR,CtILAR ORBIT PER,FORMANCE

CON FIGU RATION NO. I ST STAG E 2ND STAG E 3RD STAGE 4TH STAGE ALGOL IIA CASTOR x-259 x-258 2 ALGOL IIA ADV. CASTOR X-259 X-258 HEAT SHIELD DIAMETER - 34 IN "CU RR EN T CONFIGURAT¡ON POLAR ORBIT PMR LAUNCH

¿= I IU o t 000 f F- F- J F õu o 500 É. J :l U v.

150 200 250 300 350,_ PAYLOAD - LB5. ,|.Ñ EASTE RLY LAUNCH FROM WALLOPS ISL AN D tr I =z. I t¡J ô I000 l Il F t- J r F dl É 500 - lr. É. hi J U= v r..r o ¡F.

MOTOR STAGING 150 200 PAYLOAD - LBS.

lr!.

!.'a