a foreword to...

SPACE

Activities summarized in this report include the

major space systems, projects and studies for

which the Air Force Ballistic Missile Division is

wholly or partially responsible. Each space system and project is preceded by a condo history of administration, concept and objectives,

making the monthly progress more meaningful

in terms of total program objectives. The programs will be revised monthly to reflect major

technical and administrative changes. These

programs must be sufficiently flexible to permit

continuous and effective integration of rapidly

occurring advances in the state-of-th•-art.

2 1414.44tatot.

RICHARD D. CURTIN

Colonel, USAF

Deputy Commander, Space

AFBMD
SPACE

systems

DISCOVERER

SAMOS

MIDAS

• COMMUNICATIONS

SATELLITE

A-1 to A-9

B-1 to B-7

C-I to C-7

D-1 to D-6
A-1

WDLPM-4 106
Eyre 3. Typical DISCOVERER trajectory (above)

from larnsching Vandenberg AFB to orbit. Typical

satellite orbital path aroma the earth (right).

The DISCOVERER Program consists of the design,

development and flight testing of 29 two-stage vehicles (Figure 1), using the THOR IRBM as a first stage

booster and the AGENA vehicle, powered by the Bell

LR81 rocket engine mass as the second stage satellite. The DISCOVERER Program was established early

in 1958 by the Advanced Research Projects Agency,

with technical management responsibility assigned

to AFBMD. Prime contractor for the program is Lockheed Missile and Spoke Division. The DISCOVERER

Program will perform flight testing of the AGEMA

satellite vehicle and subsystems, and test the ground

communications and tracking network in support of

the Advanced Military Satellite Systems Programs.

Primary objectives include:

(al Flight test of the satellite vehicle airframe, propulsion, guidance and control systems, auxiliary

power supply, and telentelry, frocking and command equipment.

(b) Attaining satellite stabilization in orbit.

lc/ Obtaining satellite internal thermal environment

data.

WDLPM-4 106 A4
NOSE CONE

=COM STAG[

Weight—We

hrel MAUD

Oxidizer DMA)

ORO= WHOM Rbl14

Engles

Monk As. hoe.)

394c. kW. on. (vac)

bora Time, sec

1es1ar1 Proviskes W

Wet Weight, lbs. m)

ARENA

"A"

1,100

1.900

4.1100

1,100

XLISI4e4

15.000

277

No

2754

ANNA Iwo

1.450

3.1100

9.000

14,150

111181-116.7

15.000

277

240

Yes

265

4

LL FUEL AND

OXIDIZER

'LANK

SECTION

THOR 10011111111

Weight—Dry

Feel •

Oxidiser DOJO

ORO= WEIGHT (pass

bights

Thrust. lbs. i1.L4

spec. I.p., sat. (LW

Sera Thee, sec.

Figaro I. Photograph of two•stage DISCOVERER vehicle (left) old detailed drawl*: of AGENA, 

second siege (sight).

IG •

SEPARATION

PLANE

NITROGEN

SPHERE

HELIUM

SPHERES

ULLAGE ROCKETS

ENGINE (U0461)

IM-611

4,950

33.730

68.300

109.000

0411.31

132.000

2474

163

DM 21

5.950

33.730

611.300

1011.000

A411-311

167,000

247.11

147.5

••••

AGENA POWER 

PEW° coAsl.

REORIENT TO NOSE

APT

A-2

770 N.M.

spooss

►

tar. seo to"-

WDLPM-4 106

IHERTIAL. REFERENCE

PRoGRAMMER

HORIZON SCANNER

GAS REACnOt4 JETS

TONG COMMAND ....

compuTER

GAS JETS (ROLL)

GIMBALLED ENGINE

VELOCITY METER

.

-..p.m.......41::›

MmAND 5 ""s ....

TIME TO FIRE .....

....

COMMAND 6

VELOCITY TO

BE GAINED

1111

ci _RefsrArg. 7. 

14%14,

10a--"v

N. 4,

sit

vcrir •

Gas REACTION ./Ers

ENGINE cuT AGENA Accif.)FPEROAferEe

TRANSMITTER

Pip,* 2. Typical DISCOVERER trajectory sboM booster ellipse, reorientatim of satellite to MN WNW; Mid orbital bat erigee.
CHIN IAK

PALO ALT

00' CONTROL

cerrER)

VOICE AND

DATA CABLE

16 th ORBIT

IMPACT

AREA

WDLPM-4 106

1111111111111 . '-

45000-Rit

17th ORBIT

SEPARATION

COMMAND

CAPSULE

SEPARATION

Figure 4. Recovery sequence, showing tracking by

Alaskan and Hawaiian stations, cap.ode-front-vehicle

separation pours, end impact drat haul Ii left shows

vehicle attitude retsina' to effect capstsis re-entry

following ejection.

Testing of techniques for recovery of a capsule

ejected from the orbiting satellite.

Testing of ground support equipment and development of personnel proficiency.

•

(f) Conducting t bio-medical experiments with mice and

small primates, including injection into orbit, reentry and recovery.

Early testi confirmed vehicle flight and satellite orbit

capabilities, developed system reliability and predictability, and established ground support, tracking, and

data acqUisition requiremerds. Subsequent flights are

planned to acquire scientific data for design of

advanced military reconnaissance payload components. Typical data gathering objectives include: cosmic

and atomic radiation, magnetic field, total electron

density, auroral radiation, micrometearite measure-

. ment, Lyman alpha from space (or stars), solar radiation, and atmosphere density (drag) and composition.

A-4

A world-wide network of control, tracking, and data

acquisition stations has been established. Overall operational control is exercised by the Control Center in

Palo Alto, California. Blockhouse and launch operadons are performed at the Vandenberg Air Force Base

Control Center.

Telemetry ships are positioned as required by the specific mission of each flight Figures 2 and 3 show a

typical launch trajectory from Vandenberg Air Force

Base, and figure 6 shows schematically a typical orbit.

An additional objective of this program is the development of a controlled re-entry and recovery capability for the payload capsule (Figure 4 An impact

area has been established near the Hawaiian islands,

and a recovery force activated. Techniques have been

developed for aerial recovery by C-119 aircraft and

for sea recovery by Navy and Air Force surface vessels. The recovery phase of the program has provided

advances in re-entry vehicle technology. This Wenn°.

lion will be used in support of more advanced projects,

including the return of a manned satellite from orbit.

I" • •
DISCOVERER Wilde

No.

11.1011

Ns

Mold Dais

1959

I 1022 163 28 Febnmy

II 1018 170 13 April

III 1020 174 3 Jun*

IV 1023 179 25 June

V 1029 192 13 August

VI 1028 200 19 August

VII 1051 206 7 November

VIII 1050 212 20 November

Awaited orbit :access/ally. sksmdry received for 524 swamis after

'ifs-of.

Awaited orbit wscre.ufally. Recovery atfustle eject( or 17th orbit

was tug mama. All objectives except recover y sticceufaily achieved.

Lamb, sweat, "pare:Ws, coati awl orbital boost niccessfal. Failed

to achieve orbit beware of low paiontratwe of sweats mime.

Saws as DISCOVERER III.

All objectives awcaufally achieved except calustle recovery after

ejection we 17th orbit.

Same as DISCOVERER V.

:Wailed orbit niccarnally. Lech of 4o15-cycle Mawr preeetstai stabilisation ox orbit mai recovery.

:lush:4 orbit sacceufts117.larifiewihro prevailed AGENA *tabu

similowe t desired orbital volocity. Recovery caPutle ejected bat

sot recovered.

ereoveoar.,;.~~

WDLPM-8 106 AS
A-6

DISCOVERER PROGRAM—

MONTHLY PROGRESS

Flight Program

Two DISCOVERER vehicles were launched and

placed into orbit successfully during the month of

November. These were the first DISCOVERER flights

since August. The intervening period was used for

extensive study and testing of the recoverable capsule system.

• •

DISCOVERER VII was launched from Vandenberg

Air Force Base at 1226 PST on 7 November. On 3

November the first countdown was delayed because

of inclement weather and minor technical difficulties.

The second countdown began at 0430 on 7 November. Only one hold of 54 minutes was required to

correct a leak in the acid fill line and a malfunction

of the fuel-fill umbincal quick-disconnect. Liftoff was

normal and first stag. trajectory accurate. Azimuth

"tiers S. Lesserb ot DISCOVERER VII from Vasslimbers

Air Porte Bate ea November 7.

Syne 6. DISCOVERER served stage AGMs! "A" vehicle

ia test Thad at Sams Cam Test Base. •
WDLP81-4 106 A-7

Pigmy 7. Teleosstry main sits at

Vasieshem Air Force Bete for

DISCOVERER lights.

Piare 8. Yellers redo. Noss at Pohl, Ma a,

Ca^sforaia. Is oddities to *whin DISCOVSRER Eights, this egNipeesitt bwaintsitt

comismis 3 Led 6 to dm AGENA MOW

sup vehicle Comestand 3 is the "thee to ire"

cocrectioo aid consavvai 6 is the "velocity to

be mined" con-swims.

• ot n•' ••••

. • t

` • ! r. • •

heading was as planned and altitude achieved was

only two miles below nominal. The THOR engine

operated for 163.85 seconds lone second short of

the predicted time). AGENA engine ignition and

operation were satisfactory. Orbital status was verified by tracking stations on the first two passes and

the orbital period was established as 94.6 minutes.

Tracking operations were satisfactory, with Point

Mugu obtaining data from 44 to 400 seconds after

liftoff and Vandenberg from liftoff to 180 seconds.

Data was received by the telemetry ship from liftoff

to 690 seconds. Orbit parameters for this flight are

given in Table 1.

• • •

Telemetry received at Kodiak and Hawaii on the

first pass indicated that the 400 cycle power had

failed. Since this power is used by the satellite ohOtude control system and the D-timer, it is accepted

that the vehicle tumbled and the recovery capsule

ejection sequence was not initiated. An investigation

is being conducted.

Fronded Ached

!election Velocity. ftlisc 26.010 26,230

eerie 380 455

feriess.11•n• 104 $7

Period. We. 93.5 94.6

laclioaroa Angle, dog. 79.9 11 1.6

Ecceetricity 0.037 0.50

tilodsme, days 24 14•

'Predicted fellectiog beech

TAW 1. DISCOVERER VII Orbit Para seers
DISCOVERER VIII was launched from Vandenberg

Air Force Base pad 5 at 1125 PST on 20 November.

The countdown started at 0402 and was highly successful, with only two minor holds totaling 27 minutes

being required. The THOR trajectory was nominal

and main engine burnout occurred at 164.15 seconds. Commands 5 ("time to fire" correction) and 6

("velocity to be gained" correction) were transmitted

successfully to the second stage AGENA vehicle. A

malfunction in the accelerometer-integrator circuit,

however, prevented shutdown of the satellite engine

when the desired (command 6) velocity had been

reached and operation continued until propellant

exhaustion. The additional burning time resulted in a

terminal velocity 815 feet per second greater than

desired, and caused a more eccentric orbit than

planned. Orbital parameters for this flight are given

In Table 2.

• • •

The longer orbital period necessitated programming

of recovery capsule ejection to the fifteenth instead

of *4 seventeenth orbit to permit election within

range of the recovery forces. The ejection sequence

was initiated at approximately 1315 PST on 21

November. Telemetry data indicate that the capsule

was ejected. The telemetry ship J. E. Mann, stationed

between Alaska and Hawaii, received telemetry from

the capsule after separation from the AGENA. This

telemetry confirmed the proper operation of the

separation and spin-retro-design rocket sequence,

and jettisoning of the rocket thrust cone. Kamm°

Point tracking station also received capsule telem19smrs 9. Biomedical test corals USE-39 prior so thgradll Proile tat, sagr simaidmi orbild Capsolle is huhrunestod bower yaw :hell ad („cell.

A-13-
etry but from a higher elevation than predicted. The

recovery forces (eight C-119 aircraft and one mow

cry ship) received capsule beacon transmission but

only for a short period. No further signals were

received and the capsule was not recovered. Neither

the capsule nor the chaff were detected by the radar

aircraft. Preliminary examination of data reveals

that capsule tracking and recovery were prevented

by the following factors:

Although horizon scanner operation was excellent, satellite attitude control was lost because

of control gas depletion due to higher apogee

and greater orbital period than planned.

Satellite angle at time of capsule ejection was

probably 90 degrees instead of the programmed 60 degrees.

3. Because of higher than nominal apogee and

incorrect satellite angle, capsule impact probably occurred 700 miles south of predicted

point.

Detailed investigation of the reasons for non-recovery

is being conducted.

Predided Actual

Loom* Mies*, deg. 172 171

lefedisa Aftitv4e, 104 101

Inloam, V,Wily, ft/see 24040 26,555

In*nom Angle, deg. 0 —0.153

Isesatticky OW 0.102

Perigee, La. 104 101

AM**. Le. 350 911

Purled, aim. 93J 103.7

Dreaded MON*, dog 24 56'

'Predicted failevoing Isvach

TAW 2. DISCOVERER VIII Orbit Parameters

Technical Status

AGENA vehicles for the next five DISCOVERER flights

are at Vandenberg Air Force Base undergoing prelaunch preparations. Vehicles for the following three

Rights are at Santa Cruz. Test Bose. Two of these

were accepted by the Air Force on 17 November.

• • •

As a result of extensive radar tracking of aircraft

flights in the Vandenberg AFB and Point Mugu areas

prior to November launches probable sources of

radar interference were determined. The suspected

installations cooperated by shutting down operations

during DISCOVERER VII and VIII pre- and poetlaunch periods. Successful tracking of both flights

confirmed the effectiveness of and necessity for the

remedial measures.

Biomedical Program

Testing of a special biomedical capsule (USE-38) to

determine thermal profile test extremes under anticipated orbital conditions were completed on 25

November. Initiated on 18 November, the tests were

conducted in the High Altitude Test Simulator at the

Lockheed Sunnyvale facility. The capsule was heavily

instrumented to permit determination of the thermal

resistance between the ablative shell and the life

aril. Results of these tests should determine whether

modMcations for more heating and/or cooling of the

life cell are required.

• •

Biomedical teetimyrith , es live octant* will be

resumed 

capsule. This tat caPfgil is being P19019s1.1.9.eOrreet..

the air condilicating deficiency . which caused the

abort of the last test with a live specimen. Results of

the thermal tests of the use-38 capsule (reported in

paragraph) may require further modiRea:,

tion to theUSE-15 capsule before the live specimen . .

test is initiated.

• •

Prior to the resumption of live specimen testing,

demonstration tests will be performed on an

lifi Bill at the Aviation School of Medicine,

Broods Field, Texas. Purpose of the demonstratiOn.

veep . beiers shaw that the air regeneration problems

preciously encountered have been solved by the

current mocifficittions.

WDLPM-41 106 A-9
15.100

74.900

172.300 .

262.300

356,000

82.100

216

310

30010111—,ATLAS 101M

Weight--Wet

• Feel. IIP-1

Cleklbid %OM

. CMOS WEIGHT (Ms.)

Segleo—MA.2

Thrust Piss. vet.) Seed

Sedeleer

Spec. lop. bee. vac.) Seed

Sestehter

WDLP/4-4 106 MOM B-1

Fignro 1.

Artists' concept of SAMOS satellite.

Line dragging of complete flight

vehicle (right) and detailed view of

basic AGENA sipper stage (left).

f'

PROGRAM HISTORY

The SAMOS Program was included

in Weapon System 117L when

WS 117L was transferred to the

Advanced Research Projects Agency

early in 1958. ARPA subsequently

separated WS 1171. into the DISCOVERER, SAMOS, MIDAS programs with the SAMOS objectives

based on a visual and ferret reconnaissance system.

PROGRAM MISSION

The primary mission of the SAMOS advanced

reconnaissance system is to provide continuous

visual, electronic (and other) surveillance of the

USSR and its allied nations. Efforts include development of hardware to permit:

Determination of characteristics of enemy

electronic emissions.

Verification of known targets, detection of

unknown targets.

Location and evaluation of defenses.

Evaluation of military and industrial

strength.

Assessment of high-yield weapons damage.

Reconnoitering of troop movements.

g. Location of naval forces throughout.the

world.

!MOND STABh

Weight—WO flbs.)

Fuel (UDAVI)

Waxer 0111%4

GROSS WSW ebs.)

Ingle.

Thrust (lbs. vac(

Epees i.p. lad. um./

Sum H.. (ee.)

AMA ANNA

HA" usto

1A00 1460

1.600 3.500

000 9.1100

11,100 1030

111111•164 XU18141e.7

154000 15.000

277 277

120 240
si(4:717.4-40''

ORBITAL

BOOST

INERTIAL

PLATFORM

SW/AMUR

GIAISALLED

DONE

2 MIN

.... . .

-APOGEE OF ASCENT

ELLIPSE ($00

STATUTE MILES)

STURIMITI

DIPpagenzu.

IfrrY An/mac teR4Tir 41,40.

44.4P16 acila4C ,....11°4 WHEEL wE

CISICUtt

IaJ sm.0110".•

COAST PER

INERT

IAL PLATFORM

•

COMPUTER

REST s ars

1,4

ST4o..

6.147t err

Figure 2. Typical SAMOS trajectory, loath- to-orbit.

_ Ferret Reconnaissance...

VI/DLPP4-4 106

Fi1are 4. The Ferret spies will gather daft Pow elftwomis wwiwitmu own omfrind ly twriiwy.

. -

B-2
WDLPM-4 d.

" PHOTOGRAPH

STORE

0 0

PROCESS

. STORE

-.

ge

,

0

READOUT

•

•

STRIP VIEWED

Visual reconnaissanc•..

Store 3.

T be isitial vissel recossaivance program will use conve.tiosal photo tecistiftses with IfisoWWiC film processing amITV-type electronic image

RECORD PROCESS

stations this sh. liaR. Grand electrosict will MONvers the lige.: into photo image form, with a capability

of mole* 20 few is ksgth.

^ :

B-3

RECEIVE

Pigmy 3.

Schematic of SAMOS system

is opsrwion4 orbit. Whoa doe

tlit. is over moiriakily writer ythe Amiss syripmeet

sena os (1sformation gatb-

*riot). When it leaves usfrio.stily territory the :min

ystpstest is Hasa off and

the sessisg data is processed

(Isformation storage). When

the vehick comes witbis rasp

of a rowed receiving station,

the data will be reei-orst spots

comma for processing mid

WWWWiltill to sting agencies.

This process is continuously

repeated &grist she aid M lifetime of tbe
1

Figure 6. SAMOS concept, showing reception of commands and transmission of data

between satellite and grostnd station; and subsystem functions (schematic).

The reconnaissance equipment will be housed in the

AGENA satellite vehicle (Figure 1), which has been

flight tested in the DISCOVERER Program. During

the development phase a dual-capability visual and

ferret payload will be developed for economical

test of components. In the operational phase each

satellite vehicle will carry only the visual or the ferret

payload. The system is composed of the satellite

vehicle, ATLAS booster, launch facilities, tracking

facilities, and a communications and data processing

network.

CONCEPT

In the operational SAMOS program AGENA satellite

vehicles will be boosted into polar orbits from Vandenberg Air Force Base by Series D ATLAS missiles.

Injection into near-circular orbits (Figure 2) will be

accomplished by the AGENA vehicle rocket engine.

The satellite will be stabilized in attitude by a selfcontained guidance system using a horizon reference

scanner. As the satellite travels in an orbit essentially

fixed in space the earth rotates inside the orbit (Figure 3). As a result, each successive orbit is displaced

laterally approximately 22% degrees at the equator, permitting a single vehicle to ob

earth in a time period dependent upon the width of

the area under surveillance. Early versions will have

a useful life of ten to thirty days. Later versions will

have a useful life of one year as a design objective.

TECHNICAL HISTORY

Visual Program—The development of the

payload camera, in-flight film processor, and electronic readout were undertaken by Eastman Kodak

Company. Cameras with 6 and 36 inch focal length

lenses have been developed and the first flyable

visual reconnaissance package has been assembled.

This package indudes automatic film processing

equipment, film transport and take-up, electronic

readout, and temperature controls.

Ferret Program—The ferret payloads are

being developed in two phases. The F-1 payload was

assembled using a maximum of off-the-shelf components for early availability. The F-2 payload is being

WDLPM-4 106
..Ni.MIN

-CONFIDENTIAL

36 INCH MS

SCALE - 1:60,000

1 MILE

•

LENS - 36" FOCAL LENTa

ALTITUDE-300 STATUTE MILES

FILM - EASTMAN F5740-6

EXPOSURE- 1/100 SEC. AT F/2.8

CONTRAST RANGE - 4:1

CONTACT PRINT

ILIMSTRATING

SCALE OF IMAGE

100 MILES

17 MILES

9 X ENLARGEMENT 300 X ENLARGEMENT

Figure 7. Simsdated photography from satellite vehicle.

•

I

•

4

WDLP/4-4106 B-5
• • • e/r/Fp't

t t e:

sr -1 ;11;

SAMOS PROGRAM—motmay PROGRESS

Final assembly of the AGENA vehicle for the

first SAMOS flight was started on 9 November and

is scheduled for completion by 31 December. Operations on this vehicle at the Modification and Checkout Center are scheduled for completion by 15

February and at the Santa Cruz Test Base by 22

March. These dates are compatible with the present

launch date of 24 June. Fabrication of second stag.

components and sibassernblies for the second

SAMOS flight vehicle are proceeding on a schedule

approximately six weeks behind that of the first

vehicle. Both flights, will carry combined visual and

ferret reconnaissance system payloads.

Visual (photographic) Reconnaissance

Systems

Visual Itecommaissone System payers ere beim, dove&

aped in a Waimea mater of =maturations to attain

readout and recovery mission objectives. The destined.=

and eislective of each configaniNon is as fallen=

Readout Flights

E4 Conspenent Test Payloads

E.2 itecenersissance Payload (with 20-feet

ground essolutioni

Recovendsie Flights

El High Resolvlion .Recswersele Payload (with S400f

mound resolution/

•

Assembly of the first E-1 payload was completed at

Eastman Kodak Company during November. This

payload is scheduled to be flown in combination

with the first ferret payload on the first SAMOS

flight. The E-1 payload was subjected to 72 hours of.

continuous operation during functional tests under

simulated orbital conditions. Low contrast resoluKon

up to 88 lines per millimeter was obtained on webprocessed film. During subsequent environmental

testing, the payload was damaged extensively when

the pressure shell damping ring was improperly

secured. The payload has been diverted foi type

testing use only. The second E-1 payload, originally

scheduled for type testing use will be delivered as

the first flight article on 22 December.

• •

Fabrication of components for the first E-2 prototype

payload is proceeding on schedule. This payload

will be delivered to LMSD on 15 March. Testing and

evaluation of the engineering development model

E-2 payload are continuing. Modifications resulting

from this test will be incorporated in the first E-2

Prototype Payer.

• •

Delivery of ground support equipment for the visual

reconnaissance systems checkout area at DASD was

completed with the arrival of the vehicle power sim-

-CONFIDENTIAL

ulator on 24 November. Installation of all other

equipment for this facility was completed during

October. The primary record film processors for the

missile assembly building and the tracking and data

acquisition station at Vandenberg Air Force Base

were delivered on 23 November on schedule. installation of this equipment began on 30 November.

Technical negotiations for development of the recoverable visual reconnaissance system (E-5) ore being

conducted with the Ka Corporation. This payload

will probably use a mirror system to permit horizontal mounting of the payload in the vehicle. The camera lens will have an aperture of f/5 and a focal

length of 66 inches. A ground resolution of 5 feet

from a flight altitude of 155 nautical miles is

planned.

Ferret (electronic) Reconnaissance System

ferret NOCOSIDOWISPOI Systeme payloads also an being

developed M a ehtherse nearber of comfigurations. The

desiteatiem and purpose of each ceefiguration Is as follows

F4 RAID Test Payloads

F4 Digital General Coverage Payloads

F4 Specific Alission Payloads

The first and second deliverable F-1 prototype payloads are in modification and checkout for installation of the VHF telemetry commutator and antenna.

Subsystem checkout, prior to testing in the vehicle, is

scheduled to begin early in December. The first

payload is the flight article for the first. SAMOS

vehicle, and the second payload is a spare. Both

payloads were delivered to LMSD on 23 November,

following completion of performance compliance

tests at Airborne Instruments Laboratory (AIL). The

third F-I payload will be flown as part of the combination visual-ferret payload on the second SAMOS

flight. This payload will undergo performance compliance tests on 2 December. Formal compliance

testing will begin on 14 December, following vibrition and inhibit action tests.

•

Performance compliance testing of the Data Conversion Equipment, for use with the F4 payloads, is

continuing. Minor circuit modifications are being

made with testing of the fully assembled equipment

scheduled to be resumed early in December. This

equipment is scheduled for delivery to LMSD on 29

January, for installation in the Development Control

Center.

• •

Delivery of both F-2 payloads to LMSD has slipped

approximately 30 days because of procurement and

technical difficulties. Deliver is scheduled now for

the latter part of April.

• •

LTS414 106
'Y # 4", 4.f •

1,•••••;...ft :41#:ion

Testing of the F-2 thermal mockup payload in the

DASD high altitude temperature simulator is scheduled to be resumed in December. Results of tests

completed in September indicate that skin coating

techniques used on the mockup will permit satisfactory component operating temperatures.

• •

Fabrication and assembly of the F-2 evaluation and

command equipment are progressing satisfactorily.

This equipment is scheduled for delivery on 3 June

1960, for installation at Vandenberg Mr Force Base

and the New Boston Tracking and data acquisition

stations.

• •

The design criteria and performance specifications

for the F-3 payloads have been received by LMSD

and are undergoing technical review. Fabrication of

the F-3 payload structure is scheduled to begin in

December. The extended capabilities of the F-3

system over the F-2 system made necessary certain

modifications and redesign in portions of the vehicle

data handler. A breadboard model of this equipment

was assembled at AIL and is ready for test.

Ground Support Faellttlos

The Missile Assembly Building at Vandenberg Air

Force Base is nearing completion. Revised beneficial

occupancy dotes of 1 December for the bulk of the

building, and 15 December far the dust-free area,

have been scheduled. Beneficial occupancy of the

Point Arguello Launch Complex blockhouse began

in August, with occupancy of Launch stands 1 and 2

scheduled for January 7 and 8.

• •

Construction of the YAM tracking and data acquisition station is progressing on schedule, with

completion of various facilities scheduled on an

incremental basis through January. Rnal inspection

of the New Boston station angle tracker, command

transmitter, and UHF telemetry receiver facilities was

mode during November. The remainder of the facilities for this station are scheduled for completion

incrementally from February to September 1960.

Plans and specifications for the Ottumwa, Iowa station are complete and ready for contract advertising.

Design of support foci/Dies for this station is being

initiated. Construction is scheduled to begin in

December, with completion scheduled for February

1961.

• •

Construction of the Development Control Center

increment one will be ready for final acceptance in

December. Contract for increment two was awarded

in November, with completion scheduled for June.

Design completion for the Space Operations Control

and Data Processing Facility at Mutt Air Force Base

is scheduled for March.

WDI.PM-4 106 B-7
,

n.

Figswe

Artists' concept of MIDAS satellite.

Line drawing of complete flight

vehicle (right) and detailed view of

basic AG St# IA tipper stage (left).

SECOND STAINAINNA "S"

Ebs.) 1,450

Feel DIDAIN) 3.500

Oxidizer (MINN 9400

GROSS WEIGHT Na.) 14.550

Eagles

Thrust (lbs. sec.) 15,000

Spec Imp. (sec. rec.) 777

Sure New (sec.) 240

Figure I.

Affilti concept of MIDAS satellite.

(bottom) Linedrawings of complete

flight vehicle (left) and detailed view

of bmic AGENA *per stage (right).

•

SOLAR

A RRAY

DESTRUCT

CHARGE

ATA LINK NOSE CONE TM EXI EQUIP OXIDIZE FUEL

ANTENNA SEPARATION ANTENNA COMPT. TANK TANK

•

S./

SOOSTER.--AlLAS KIM

We1/401--Wet 15,100

, Feel, W-1 74,900

Oxidizer (LOX) 172,300

GROSS WEIGHT (bs.) 262,300

Thrust (lbs. rec.) Rest 356,000

Smarter 112,100

Spec. Imp. (sec. roc.) Soot 216

Wakes 310

C-1
CONDMONS:

2,000 sm. altitude

vehicles hi polar orbit

1800

Reecleut stedees

Enfloesi

Nee* Mem*

North Pee&

60°

Figure 4.

Orbiting satellites detect infrared radiations

released by Soviet ICBM 'winches. Data tel-

*metered nuianianeomsly to MIDAS Control

Center via far north realms stations. Decoded

data reveals EXACT number of missiles

launched and APPROXIMATE launch location, direction of travel and burning characteristics. Map abOve shows percentage of

detection probabilities over USSR.

•

TECHNICAL HISTORY

The MIDAS infrared reconnaissance payload

will be engineered to use a standard boostersatellite launch vehicle configuration. This

configuration consists of a "D" Series ATLAS

missile as the first stage, and the AGENA

vehicle, powered by a Bell-Aircraft rocket engine, as the second, orbiting stage (Figure 1).

Refinements to the AGENA vehide will be

made as a result of the DISCOVERER flight

test program. The first flight article infrared

payload has been assembled and Installed on

an AGENA vehicle, and checkout oPenifians

initiated. A solar auxiliary power unit has

been developed and fabricated for installation on the third flight. The third major component of the payload, the communications

package, also has been designed, fabricated, and tested. The total payload weight

has been determined to be 1,014 pounds.

The ATLAS/AGENA configuration with single

restart capability and large propellant tanks

can place a payload of 1,500 pounds on

2,000 nautical mile altitude polar orbit (see

Figure 21. Only the first two R&D flight tests

will use the single capacity AGENA vehide.

WDLPA1-4 106 •. •
ENTER ELLIPTICAL ORBIT

PERIGEE OF

COAST ELLIPSE

AGENA

1ST. BURN

---\ 150 N. MI.

VERNIER/SEPARATION &

". SHORT COAST

SUSTAINER

•

BOOST

Pipes 2.

Leancb-to-orbit trajectory. Optinsans ATLAS boost, voided

by radio-inertial system. AGENA Ismer (coast, bens, coat,

second Mrs) provides amitosis refrINCe. Also governs, velemsty nsagninwde wad direction by Menial vigour.* system monitored by horizon scanner. Orbital MallIkiall by

ruction wheel aid gas jets.

PROGRAM HISTORY

The MIDAS Program was included in Weapon

System 1171 when WS 1171 was transferred

to the Advanced Research Projects Agency

early in 1959. ARPA subsequently separated

WS 1171 into the DISCOVERER, SAMOS and

MIDAS Programs, with the MIDAS objectives

based on an infrared reconnaissance system.

The MIDAS (Missile Defense Alarm System)

Program is directed by ARPA Order No. 38,

dated 5 November 1958. Development activities will lead to the first of a ten Mght R&D

program in January 1960, with a reliable

operational system scheduled for 1962.

WDLPM-4 106

180° COAST

AGENA 2 ND. BURN

ENTER

CIRCULAR ORBIT

OFtIENTATION

2000 N. MI.
low.

... ...

INOIVIMIAL

411111111001111

CUMIN agar

TV

Figure S.

Simplified version ofgonad presentation system (left) for display of

infrared reconnaissance data. The

data is displayed on a TV monitor

with a map overlay. The chart below

shows data flow from the reading

stations to decision-making agencies.

Tim MIDAS Control Center, or other

using agencies having a correlated

ground stabilized display, can determine when an ached attack has been

launched. The decision to cosintaattach is made by the President, with

all elected agencies reacting as preplanned.

ICBM ATTACK SITUATION

CONCEPT

The MIDAS Program system is designed to

provide continuous infrared reconnaissance

of the Soviet Union. Surveillance will be conducted by the accurate positioning of eight

satellite vehicles in orbit (Figure 3). The area

under surveillance must be in line-of-sight

view of the scanning satellite. The satellite

operating altitude is determined by system

infrared 'scanning design. Minion capabilities are shown in Figure 4. The system is

designed to accomplish instantaneous readout of acquired data by at least one of

three strategically located readout stations.

The readout stations transmit the data

directly to the MIDAS Control Center and

other display centers, where it is processed,

displayed, and evaluated (Figure 5). If an

attack is determined to be underway, the

intelligence is communicated to a central

Department of Defense Command Post for

relay to the President and all national retaliatory and defense agencies. The decision

to counterattack is made by the President

of the United States, with retaliatory forces

reacting only in response to the Presidential

decision.

—CTRIFFBENTIALNDIpm4 106

t! '
-1friNiSr4,00t4t,!t.....4 fi.,!,T.4 72: T,,rt,t7Cel,ret

. . -

AIM C-5

to; i

•

1111111111r

MIDAS PROGRAM•MONTRLY PROGRESS

Program Administration

Objectives of the MIDAS flight program have been

revised to attain polar orbits at 2,000 nautical mile

altitude commencing with flight 3. The 1,000 nautical mile orbit _flight has been deleted from the

program. Rights 1 and 2, scheduled for February

and April 1960, remain unchanged.

Technical States

The second stage AGENA vehide for the first MIDAS

flight will undergo a second hot firing at Santa Cruz

Test Base on 2 December. Shipment to AMR is scheduled for 6 December, following Air Force acceptance. The AGENA vehicle for the second MIDAS

flight is scheduled for deliver? to Santa Cruz Test

Base on 12 December.

• •

An integrated test of the ground-space communications data link van and the Aerojet-General scanner

unit has been completed. The test verified compatibility and generated tapes for use in future training

and checkout.

• •

LMSD representatives have evaluated the General

Electric proposal for an operational type ground

presentation unit. This proposal was the most satisfactory in response to a preliminary work statement.

As a result of the evaluation, a detailed work statement is being developed.

• •

A test program conducted by Infrared Industries,

Inc., indicates that the probability of solar damage

to the orbiting infrared scanner is negligible.

• •

The pneumatic aftitude and orbital control systems

configuration for flight 3 has been established. This

system controls attitude variances which exceed the

Spore 6. Second Aerojet-Generai

Scanner snit shows wasted is High

Ahead. Tempovasse Sinesdator

prior to ten.

WDLP14-4 106
limits of the reaction wheel system. The bask electronic module design concepts for the orbital control

gas system haw also been determined. Mechanization of the complete electronia will follow determination of the number of reaction lets to be used.

• •

The vibration PAM/FM telemeter, planned for ascent

instrumentation, will be converted to a high-response

PAM/FM telemeter for Mghts 3, 4 and 5. 'this will

allow the unit to be used in orbit for radiometric

data in addition to the ascent vibration functions.

• •

Launch complex 36, Patrick Air Force Base. Benstidal occupancy dates for the blockhouse and launch

stand are scheduled for 18 May and 15 July respectively. Final plans for the propellant loading system

and the service towers were reviewed on 10.11

November. Plans and specifications will be released

for construction on 1 December.

• •

Tracking and Data Acquisition Stations. All foundation work for the technical facilities at Donnelly Flats,

Alaska, has been completed, and the administration

and data acquisition building, and the power plant,

have been enclosed. The contractor closed down

operations on 1 November for the winter and will

resume work on or before 1 April. Completion of

the various facilities is on schedule, and will be

completed on an incremental basis between June

Ai 7. CoAwe. for we with the

ix Ai smatter :licit At the All

rile Rene, Missile Artaxily

Building:

C4
Pip.. 8. (kit) AGENA semi' stage

vehicle (No. 1007), have subjected to beerier fep/vajox allproemial...

This maide it se for itutallatiox

or the tenni MIDAS fiighe vehicle.

Figaro 9. (above) Sword stage of first MIDAS

Iligk rebid* stadergohtg vertical weighing

at the coselatios of the tvoslificatioa sad thechoes proceslare. Nozzle of AGENA solidi nem is sysible et bottom of photo- groh, aitrogn aad believe sphere: oat either side.

,,

•

Figaro 20. (left) Initial her-Brixg of AGENA mood stage vehicle (for first MIDAS flight) at Santa Crux Test Btu..

. WDLPM4 los • •

•
IR SCANNER

COMMUNICATION ANTENNA

TRACKING/TELEMETRY

ANTENNAS (4)

NOZZLES

SUN TRACKERS

PHOTOVOLTAIC CELLS

Fiore 1. Cotamaxicatiotu satellite (STEER Phase)

rebid*, showing aatettia ma solar paddle *rime-

'ion. Small dements show opetatiostal coarept of

STEER phase.

WDLPM-4 106

Communications

Satellite

The Communications Satellite Program will investigate, in three phases of increasing complexity, the

feasibility of using synchronously spaced satellites as

instantaneous repeaters for radio communications.

Under ARM Order No. 54 as amended, AFBMD is

responsible for design, development and test of the

complete system, induding launch, satellite tracking

and control, and necessary support facilities and

units. Wright Air Development Center is responsible

for the development of aircraft communications

equipment for all three phases. Responsibility for

satellite and ground station communications equipment is assigned to WADC (first phase) and to the

Army Signal Research and Development Laboratory

(second and third phases). The three phases of the

program have been designated STEER, TACKLE and

DECREE. The description and objectives of each

phase are as follows;

STEER (Figure 1)

This four-Right test phase will use ATLAS/AGENA

vehicles to inject satellites into polar orbits with six

hour orbital periods. This phase stresses earliest possible availability consistent with program objectives.
AGENA SECOND STAGE

MAX. XLR-Ba81-7 engine

90'

69/

ATLASS BOOSTER

Thrust (lbs at sea level)

Main Engines (2)

Sustainer Engine (1)

Vernier Engines (2)

Total at lift-off

309,000

57,000

2,000

368,000

Pip,. 2. Night vehicle /or STEER owl TACKLE

STEER objectives include:

Provision of a single channel, two-way voice

communication repeater between ground stations in

the United States and airborne strike-forces of the

Strategic Air Command flying alert missions in northern latitudes.

Development of engineering concepts and

equipment, and furnish test support data for TACKLE

and DECREE phases.

c. Investigation of the effects of vacuum and

radiation environment on satellite components over

an extended time period.

The four vehicles are to be launched from the Pacific

Missile Range starting in March 1961. The vehicle

(figure 2) consists of an ATLAS booster and a modified AGENA "B" second stage (double capacity

propellant tanks and single restart capability). The

payload will be placed into a circular orbit (figure 3)

with a period of one-fourth of a sidereal day. The

5,600 nautical mile apogee of the transfer ellipse

apogee is reached during a coast phase following

first shutdown of the AGENA propulsion system.

When apogee is reached the AGENA engine will be

reignited to attain orbital velocity. After AGENA

shutdown the final stage vehicle will be separated

from the AGENA. An attitude control system will

then orient the payload antennas toward the earth

and solar cell paddles toward the sun to permit

communications system operation.

- •

COMMUNICATIONS SUBSYSTEM—The three elements of this subsystem are the ground station, the

satellite repeater, and the aircraft communications

equipment. The 10 KW ground station transmitter,

operating in the lower portion of the UHF spectrum,

will use an antenna large enough to provide maximum reliability in the possible presence of interfering

signals. Initial test antennas will be 40 and 60-foot

parabolic types. These will be replaced with hardened or semi-hardened antennas later in the program

to provide compatibility with the SAC hardened control centers. The receiver will use the same antenna.

The ground antennas will have a tracking capability

to keep the antenna properly oriented toward the

satellite. Both simple frequency modulation (with a

deviation ratio of about 4) and pseudo-noise modulation (spread spectrum) will be tested during this

phase.

Since the function of the satellite repeater is the realtime relay of messages, separate reception and

transmission frequencies (10 to 15 megacycles apart)

are to be used. Satellite transmissions will be at a

40-watt r-f level, using FM modulation and a 10-

decibal gain antenna.

When properly oriented toward the earth, the satellite antenna coverage is approximately the same as

the angle subtended by the earth at a 5,600 nauticalmile altitude. Coded control circuitry in the satellite

will permit the repeater to be turned on or off to

avoid undue power consumption when not in use.

Except for the final r-f power stages, the entire

repeater will be transistorized. Low noise preamplifiers will . be used because of the low-level signal to

be expected from the aircraft transmitters.

Aircraft communications enuipinent will make maximum use of UHF equipment presently installed on SAC
Rpm 3. Schematic of STEER Owe flight make,

tar, mod orbits showing wow* of oimwoiklmi

moos.

TIME ALTITUDE VELOCITY

(SEC) (FEET) (FEET/SEC)

BOOSTER SEPARATION 139 208,000

SUSTAINER SEPARATION 295 • 566,000 21,000

'STAGE II IGNITION 297 567,000 21,000

STAGE II 1st CUTOFF 395 570,000 30,800

STAGE II 2ND IGNMON 6649 34 x 106 12,100

(ORBIT INJECTION PHASE)

9,400

STAGE II FINAL CUTOFF 6667 34 x 10' 16,000

SATELLITE SEPARATION 6677 34 x 10' 16,000

aircraft. Minor modifications will provide an adequate

signal-to-noise ratio from the satellite transmitter at

a maximum range of 8,000 nautical miles. Three different approaches to aircraft-to-satellite transmissions will be tested. The first, a modification of the

aircraft ARC-34 equipment to permit FM modulation

capability, is expected to be marginal in performance. To provide better performance an FM receiver

transmitter is being developed which would replace

the ARC-34 unit, providing 150 watts of transmitter

power. A 1 KW radio transmitter is also being developed to replace one of the. AN/ALT-6 trdnsmitten.

This unit will be capable of the secondary function of

providing a certain amount of jamming power in the

UHF band when not in use as a communications

transmitter. Should spread-spectrum modtdcdion be

required to overcome interference or jamming on

the aircraft-to-satellite link, compatible modifications

would be made to aircraft equipment.

GROUND SUPPORT FACIUTIES — Ground tracking

and data handling capability is reqUired to: (a) Verify that the satellite has been injected into orbit,

(b) provide data on performance of the Anal stage

vehicle in orbit, and Id Provide sufficient control to

permit synchronization of the satellite position in

relation to other satellites in the total system. IMPSligation is being mad, of the possibility of using

SAMOS and MIDAS Program ground support facilities. Ground stations at Offutt Air Force Base and at

each of the three numbered SAC Air Force units will

provide the capability for two-way communications

via the satellite. Each of these stations will be able

to compute the precise position of satellites in the

operational system.

TACKLE

The four flights in this phase also are programmed

for use of ATLAS/AGENA vehicles to attain polar

orbits having six-hour orbital periods. TACKLE

objectives are based on the use of the STEER phase

test vehicle to permit early Moist testing of microwave components and other equipment required for

DECREE phase. Flight trajectory and orbital parameters are the same as for the STEER phase.

DECREE

The seven flight tests in this phase will provide

the R&D effort for demonstrating the feasibility of

a global communications system, using precisely

spaced "hovering" satellite which, essentially, have

orbital periods of 24 hours.

WDLPM-4 108 .

. . • MEP .•

-GONFIDENTIAt D-3
COMMUNICATIONS SATELLITE SYSTEM.

MONTHLY PROGRESS

Because of the long-life requirements of the Communications Satellite, great emphasis is being placed

on reliability in the design and fabrication of systems, circuits and componenb. Since many of the

required components have not yet demonstrated the

necessary useful life expectancy, extensive design

analysis, supplemented by laboratory studies will be

necessary. Redundancy will probably be one of the

primary means of attaining the required reliability.

However, the five-months postponement of the first

launch date, will permit increased reliability to be

incorporated into initial design efforts. The present

reliability requirement is for a life expectancy of one

year in orbit.

• •

A new launch schedule has been formulated for Point

Arguello Complex #1 to integrate the Communications Satellite launchings with the scheduled SAMOS

and MIDAS flights. This new schedule provides an

additional Ave months to be used for design efforts

toward increasing reliability, development of orbital

correction capability, and studies of optimum antijamming techniques. Action has been initiated to

obtain approval for location of a ground communication station at Mead, Nebraska.

PROJECT STAIR

Program Administration

A contract for the satellite vehicle has been negotiated with the General Electric Company, Missile

and Space Vehicle Department. The contract probably will be signed during December. A contract for

the communications subsystem was signed on 29

October with the Bendix Corporation, Systems Division. Space Technology Laboratories has completed

a study of the Bendix proposal for the subsystem.

Results of this study will be presented to WADC and

Bendix in December. If the required funds are

released immediately by ARM, the contracts for the

first and second stages will be negotiated in December. A work statement for the ATLAS first stage

booster has been prepared and is In final coordination before being submitted to Convair, Astronautics

Division. The Lockheed proposal for the second stage

vehicle was received on 2 November, and has been

evaluated. A briefing for the Ground Tracking and

Data Acquisition contract is scheduled for early in

January 1960.

• •

A Project STEER Development Plan, reflecting revised

flight =Nadu* and funding 'requirements, is being

prepared. This Plan will be submitted to ARPA in

January 1960.

• •

A special ten-man team from the WADC Eledronia

Technology Laboratory has been formed to monitor

the Bendix effort to insure that required component

reliability is attained in the communication subsystem.

Technkaf Progress

Technical direction meetings held during October

emphasized satellite vehicle and communications

subsystem development plans, including the new program schedules and orbit correction equipment. The

following agreements were reached,

The satellite vehicle contractor will submit

revised schedules immediately, and the program plan

by 1 December.

The communications subsystem contractor will

submit a program plan for FM dear transmission and

anti-jamming equipment by 1 December. The feasibility of combining the FM and anti-jamming equipment will be studied. Simultaneous, as well as

alternate, opeiatien of FM and anti-jamming equipment will be investigated. A fail-safe feature would

select the FM receiver if the anti-jamming receiver

should fail.

• •

The first complete satellite flight vehicle will be available by 15 December 1960. A flight proofing test

will be conducted during January and the vehicle

will be delivered in February 1961.

• •

Trajectory studies, using the most recent performance

figures for the ATLAS/AGENA "V' propulsion systems, indicate the maximum allowable final stage

vehicle weight to be 600 pounds. General Electric

Company is evaluating the estimated weights of

components based on this maximum vehicle weight.

• •

During a reliability discussion on 17 November,

ARPA expressed concern about the reliability of the

satellite-borne pseudo-noise (spread-spectrum) equipment. As a result, Magnavox will redesign this

system extensively to incorporate the latest circuit

reliability techniques. lisnsdix engineers are assisting

to insure that reliability prediction calculations are

consistent with those being used by Warm and

Space Technology Laboratories.

• •

On 13 November, a meeting was held in Washington, D.C. to discuss the communications frequency

to be used by Project STEEL WADC/Bendix were

directed by IbrUSAF (APOAC) to study the effects

of Axing a frequency in the vicinity of 400 MC

100111111E -CONFERRAL WDLPM-4 106
rs .10 11,VC

instead of the requested 225-280 MC range. The

study revealed that the most severe effect would be

to the satellite-borne equipment. The presently

planned transistorized signal amplifier stage would

have to be replaced with a hard tube stage, resulting in decreased reliability. Also, the aircraft equipment would require tube development to obtain the

power output specified within ski and weight

limitations.

PROJECT TACKLE

It has been recommended that this Project be incorporated into Project DECREE.

PROJECT DECREE

Program Administration

Available payload space on the last Huse CENTAUR

program R&D nights will be used by Project DECREE

for preliminary development of the satellite vehicle

and communications payload. This will result in

decreased booster costs and provide the additional

advantage of using the satellite vehicle in its

required orbit.

• •

The DECREE Development Plan, currently being prepared, is scheduled for presentation to ARPA during

February 1960. A preliminary spedfication for the

-CONFIDENTIAL

DECREE system was presented to ARPA on 10

November. This speacation proposed the use of a

2 KMC communication system as capable of providing greater over-all system reliability than an 8 KMC

system. However, in accordance with an ARPA memorandum to AFBMD, dated 13 November, design is

being oriented toward use of 8 KMC transmission to

the satellite and 2 KMC reception, from the satellite.

Capabilities of both X-band and S-band high power

amplifiers, and other X-band components, are being

investigated by USASRDL

Technical Progress

An itemization has been prepared of the supporting

research and development which must receive early

funding. An investigation is being made to determine

if similar developments are in existence prior to submitting this list for ARPA approval.

• •

General Eiedric Company has been directed to

obtain more detail on possible anti-jam operation,

particularly in defining problems of automatic frequency control and pseudo-noise generation.

• •

A contract has been awarded RCA for a 24-month

effort in the development of a highly reliable penal

tube, Type 5876, intended to be used as a frequency

multiplier driver.

WDLPM-4 106.
L SPACE

L

projects

ABLE E-1 to E-9

TRANSIT F-1 to F-5

TIROS G-1 to G-3

COURIER H-1 to H-3

AGENA J-1 to J-3

ABLE-STAR IC-1 to K-3

MERCURY L-1 to L-6

609A M-1 to M-3
Figs., I. ABLE-3 fligbe sea tebide being larnebei from Askonie Mail* Reno. Din gesuional drawing. (left) of /ors-seage ABLE-3 Mak. •

•

WDLPM-4 106 ONE E-1
,

s ' . • r'.1 •s

Sion 2. ABLE-3 thinf stage ood paylirai with solar paddles is fslly oxkoloi poaioo.

The ABLE Program consists of three flights with primary missions as follows:

ABLE-3: This initial flight placed an advanced

scientific observatory satellite in an extremely elliptical geocentric orbit about the earth.The payload contained instrumentation for thirteen experiments, and

solar power sources for long term electrical power supply. This satellite is obtaining unique and original scientific data as well as testing components for the

ABLE-4 program.

ABLE-4 ATLAS: The second flight will place a 371

pound instrumented payload in orbit around the moon.

This flight will investigate the space environment in the

vicinity of the moon, and provide crude television

images of the far side of the moon. The large percentage of the equipment used is identical to that installed

in the ABLE-4 deep space probe and Lunar satellite

vehicle.

Pik 3. Drools: of morosely ollipoical ABLE-3

E-2

+,t4

WDLPM-4 106
ION CHAMBER

GEIGERHULLER

COUNTER

TEMP RANGE

REF

SEARCH COIL

MAGNETOMETER

LOGIC

(UNDERNEATH)

ASPECT

INDICATOR

CONVERTERS

NO.211NO.3

40 SECOND ACCELEROMETER

TIMER BATTERIES

INJECTION ROCKET

HEAT SINK

SW TRANSMITTER

BATTERIES

COMMAND BOX

DIGITAL

DECODER

TELEVISION

LOGIC CIRCUIT

COMMAND RECEIVER

SOLAR CELL

CURRENT MONITOR

FLUX GATE

MAGNETOMETER

SUBCARRIER OSCILLATOR

AND

MULTIPLEXER "B`

BATTERIES

SUBCARRIER OSCILLATOR

AND

MULTIPLEXER se

DIGITAL

TELEMETRY

PROPORTIONAL

COUNTER

(CHICAGO)

STATIC

CONVERTER

10B MC

TRANSMITTER

THOR — ABLE III PAYLOAD (TOP VIEW)

Pigott 4. ABLE -3 payload (sop

c. ABLE-4 THOR is a deep space probe which will

be launched into a solar orbit intermediate between

the orbits of earth and Venus. At its closest approach

to the sun, the probe will pass near the orbit of Venus,

returning to intersect the orbit of earth at its greatest

distance from the sun. The vehicle consists of a THOR

first stage, AGC 10-101 liquid fueled, guided second

stage, and ABL248A-3 solid-fuel third stage. This

vehicle will place an 88 pound payload into space

flight. The payload contains no retro-rocket, and no

attempt will be made to intercept Venus. The design

of the payload components is the same as those which

hove been proven on the ABLE-3 satellite.

The ABLE-STAR Program will develop a versatile and

efficient upper stage for use with varied booster/vehicle combinations. This stage will have basic design

features proven in the dependable AJ10-101 stage

used on the THOR/ABLE vehicles.

•••4 eit

• WDLP24.-4 106
•• •

.s•

• ! •

E

•••1

id

r4

D4 ....

....•1

AI rd

•

COm

0•••

4

CO ,13r, O•

v.,

•

r4 .s.

CA

r,3

•

i

4 n CO

n. 38

4

0% co co

•

1 as

1

Q es

u.4

4 so

,

•

•

•

a 9Q

N

r4

CO

•

0,

N

31:12S,'

.4

•

0•

r4

0

•

04

N

•

*0

CO

a

N

at

VI

0"

oe

2

%. •

0

at 0

. 14

14 4.1

•

ft4

00 CO

1.11

1

A

43

ini m4

VI

A.

•

1-4

NI

VII

r4

CT •

en*0

ft

4

1•••

PI; In

41.

tell

1•4

•

11.•

1111

C)

Ps C1

en

m

4•1

••

•

COat

4' •

en

CA

•

910

•••

CO

•

.4CO

EnCo

• Mm

gs

N•

411

•

N

PA

•

.., 0•

•

to

co co

h

Ch

• N

re

14NI AI

0

k op g

CO CON

Al

o

CO

.

on1.•

VI

.

0. Ps NCO

VS

go

on

•

en111

1,1

CPn

c•

••1

6.4

.

PP

•

8i:14'.4!

IA

004

CO•

e%N

o

Tn

e r

0

10

N

•

I *0P.

•

CO

ch

N

P

...

'0

•

A

40

•

CA

/•n

ot

•

NN

0

0

4

CA

C*1 N

•

COo'd

•

•••

CC

.

C•

Ps

•

ap

hn

0 0

CO

CA

d

N.1

*0

OLP

N

1

k 4.1

0

A4

0

•

m

*0

r40

•

C4 mo e-t

OD

13% O.•

Al

4

.4

Ps

pi

cr••

0

rd

en

^IAel•

CO

CA

en

Pn

0

•

el Ch

IACO•

0%00

14

•

0

r4

e•

• WI

r4

••1

•

4)

en12

• WI

r4

..4

to

•

1.•

oo

le

•

tY, In

es

VI

•

el

IA

1•••

•

est

rtet•

••

r4

CO

•

CO

r... 0

en1,11

moon

PI

N

•

co

CAN•

m 00

m

N•

V

0

10

+

1.4 4.1

g

a

on

a

0

r,V,338

*0

e-o

•

1•4

coch

0

•

r•

va

"N•

IA

co

as

4,

III

NI

CA

en

.4

•

S

Cli

•••

. •

S

0

•

esd

NO.

11; PI

N•

eh?•• 0

4'

r4

•

le

•

4

'0•

I.

0

..4• N

en

101•

:A

O•

111

••fek

h

•

Pn

••0

44;33°

Ts 0

••

PA

r4• CO

• 44,,,,

VI

In

•

3

N4..,

Irt 4.,•

0

Ir3

N

co

P.: m4. AI

a,

r4

14

VI Irt

irl

CAin

O

en

cv

41: in

CO

r4

VI

,•;

r•I

144 8 u 0 •••4 4

41 IC

4•. C1 14 •••1

DI fo• b.

0

r4

0

4.4

enI•1

*0

P.•

En*0

0 0

en.0•

A

VI of•

coCP*•

OD

N

0

•

330.00.00 ••

Ch•

en.4

a%•

en14

0%-1• • O

,0

..

ION

0;

.00

!

f

rn•

4

rt

,...

en•

4

in

...

00000000.,.40 000 0000

•

O•

•

at

•

co

•

ch

88§§0000 • • • •

00

•

i

13

U

"I

it

AI

P4

U

4

g i

14j

U

k

Pe

•

r4

0

AI

A

ns

•

AP

0

i

04

rd

k

u

4

-1 3 .'3:6 2

o

UY

U

he

a

14

u

.4

os

e

U

U

Se

a

.4104101k ' 3 :g

.4

u

44

.0

co&I

0

Se

a

PI

U

4

mo

•

u

0

I. Pe

3414 4044

r4

t.) 4

2

U

•

t•Pe

14.1

r•1 U

Er34•

u 14 4 Pe

.4

1.1

a

4

34

V•

AP

•

•

sa

la,

.•I

g4.1

0

43

. .

t .

•Ve

co

•

I

f

.

idke

•

e

la

..

1.1

”I4

•

roA

06 

co N

464 $4

C13

is

u

•••1

u

0

••I • > • 0 u

CO

r•I ...,

13

%4

•

101 AI

64

CI

14

let 4.1

d44)0

te

r•I01

is

•

P4 al

4.

Al 0 14

111

1.4

g

PP 00

•

s• a

r4 a

4 .0 4 4.1 41 g

14 4

04 4

••••►..

1111111111111CONFIC,7:nar

3 •

0% • 4.1 0

..4 •

S th

.-•.4.• 10• O.

$4

4

•

5

id0

•

0% se 4.1 ON 01 00 g it

g X

0

..4

•

U .0

P. e

.g ••n

,I

1 it 4 A

•

4

a

/

0

4.4 41

44 le

O e

A

o-400

WDLP1 -4 106
Improvements being made include: (a) increased propellant capability; (b) a muhiple restart capability;

and (c) a full-time attitude control system (to operate

during coast periods as well as powered flight). These

improvements will permit a two-stage THOR/ABLESTAR vehicle to attain weight/altitude performance

equal to that of the three-stage THOR/ABLE vehicles.

This will provide increased reliability and accuracy.

This stage will be suitable for mating to THOR, ATLAS

or TITAN space boosters and will also accept the

ABL24I3A-3 third stage, if needed.

*AM t.1,01,

Nesse 1 Liao drawls: of ABLE-4 ATLAS Big& tea which. Moser show "skids cm

lostscb id. Mick boosto secookstags awe slim/ma* kora* kirks, /light

Aria: mains=

WDLPM-4 106
I 73 .11 ;Pk I "a a

• Ir V " 4

liiptre 6. Galway visa,

 of ABLE-4 THOR Akght test whisk (WA Drawksg (right) :how: or gies coaceptioa of A8124 THOR Payload as orbit alma the

E-6

WDLI44-4 106
ABLE PROJECTS.-.—MONTHLY PROGRISS

A1111-4 ATLAS

ABLE-4 ATLAS was launched on schedule from the

Atlantic Missile Range at 0226 EST, 26 November.

At approximately T plus 48 seconds, the second stage

Advanced Guidance System transponder ceased

operation, and parts were seen to break away from

the upper stages. Four to live seconds seconds later,

additional parts were observed falling away. At

that time the rate gyros in the ATLAS recorded a

disturbance of about three seconds duration and the

axial accelerometer indicated a slight decrease in

vehicle acceleration and a change in attitude for

approximately 4 seconds. These indications tend to

confirm the breakup of the upper stages. Second

stage telemetry equipment ceased to operate at

approximately 104 seconds. The ATLAS continued

until shut off by its guidance system after successful

completion of the planned first stage portion of the

Right. Two factors indicate that second stage ignition occurred at the planned time: (a) the ATLAS

axial accelerometer indicated a decrease in acceleration which could have been caused by second

stage engine blast impinging on the ATLAS, and

Ibl the ATLAS control system gyros recorded conditions similar to those observed on THOR boosted

Rights at the time of second stage ignition. The nose

faking, payload and third stage probably broke

away from the second stage at T plus 48 seconds

causing the disturbances recorded by the first stage

equipment. Indications are that this failure was

caused by one of the following three malfunctions,

in order of greatest probability:

The nose fairing parted due to structural or

latch failure.

Acddental ignition of the payload propulsion

system created excessive pressure inside the

nose faking.

3. The payload broke away from the third stages

or, payload and third stage broke away from

the second stage, due to structural failure.

Recovery of some payload parts in the surf off Cape

Canaveral confirm that the payload was detached

from the vehicle after approximately 48 seconds.

Studies are being continued and it is anticipated

that a final evaluation will be available for inclusion

in next month's report.

• •

The ABLE-4 ATLAS project was the first use of an

ATLAS ballistic missile with stages as booster for a

space payload. The ATLAS booster performance was •

Pigmy 7. ABLE-4 ATLAS Miami (solar paddles

tot issialk1).

•

WDLP144 106 111110111.It . ::f.., .. E.7
WDLPM4 106

tüllIMENTIAE i

E-8

completely as planned. This flight lends confidence

to the reliability of ATLAS performance in future

flights. The ATLAS is programmed for use in the

Advanced Military Satellite Programs and the Project Mercury (Man-in-Space) effort.

• •

The ABLE-4 ATLAS trajectory from Cape Canaveral

to the vicinity of the moon was selected to achieve

the tightest circular satellite orbit around the moon

consistent with the highest probability of success.

The final burnout conditions were established at an

energy level corresponding to an inertial velocity of

34,552 feet per second at a total distance from the

center of the earth of 23,290,000 feet.

ABLE-4 THOR

Program AdmMistratioa

The Detailed Test Objectives trajectory was made

available during the month. However, changes were

made necessary late in the month by the change in

second stage. Launch is scheduled for 15 December,

between 0715 and . 0745 EST. Free flight and powPiga.. 9. Cbrckiag solar cell paddle installation or ABL&4

ATLAS poised.

More 8. ABL1I-4 ATLAS Payload asseaddy for third bags

(solar paddles instalksi and wailed).

grad flight trajectories for four days subsequent to

15 December also have been computed. The payload weight estimate at the present time is 90.4

pounds.

• •

The ABLE-4 THOR payload is a deep-space probe.

A communication range of 60 million miles has been

set as the goal for the tracking, telemetry, and command subsystem for this experiment. In order to

achieve this range objective, more sensitive receivers

are being installed in this payload, and the Manchester station equipment is being very carefully

reworked to achieve full transmitter power and best

possible receiver sensitivity. The transmitted commands from the Manchester station, at 60 million

miles range, are almost exactly at the level required

for the payload receiver threshold in the narrowband

position. It can be concluded from this that the

command subsystem is very likely to function properly to at least 30 million miles, and may function to

60 million miles.
V""'srirt

Testlag Program

During the months of October and November 191

environmental acceptance tests, type tests, R&D tests,

and parts evaluations were performed, with only

five unsafisfadory performances. All five of these

were of a minor nature and have been corrected

satisfactorily.

Ground Support Program

An improved set of guidance and telemetry antennas

was fabricated and checked out in Los Angeles. A

high temperature material (fiber glass) with a high

temperature resin was used. This material will not

begin to ablate until temperatures of 1200.1400

degrees F. are reached. Ablation on the other

antenna she began at temperatures between 500

and 600 degrees F.

-CONFIDENTIAL

operable at full power because of unsatisfactory

transmitter tubes. New tubes are being delivered

which are expected to meet specific performance

requirements.

• •

HAWAII—Effective 12 October, the full time maintenance of the three diesel generators was assumed

by the Pacific Missile Range. The UHF 10KW transmitter is not operable at full power because of

unsatisfactory transmitting tubes. Because of feed

difficulties on the 60-foot antenna, arrangements

have been made to permit the use of the bidet

antenna for hansmbsion purposes, if necessary.

• •

SINGAPORE—This station is ready for the ABLE-4

THOR launch.

ABLI-3

The Space Technology Laboratories Space Physics

Data Library is now in operation. Already stored

within the Library are 600 magnetic tapes from

ABLE-3, as well as tapes from Pioneer I and Pioneer

11 (ABLE-1). The Library is preparing a catalog of

ABLE-3 trajectories. The bulk of the ABLE-3 digital

data had been reproduced by 9 November. After

having been checked and cataloged, the applicable

data was mailed to Stanford University, University

of Chicago, or the University of Minnesota. The second osdilogram reading table was installed in the

Library during November.

* •

MANCHESTER—During the week of 9 November,

the linear polarised UHF feed was shipped to Manchester. The work of mounting the Rantec antenna

assembly on the roof of the Telescope Control Build-

- big was completed. Use of this location will permit

improved low elevation angle visibility towards the

west. Following modifications to the 10KW trans-

_

mitter, accomplished in October, satisfactory operation has been obtained several times for extended

periods. At present, the 10KW transmitter is not

E-8
Fiore 1. TRANSIT IA being kursulted from Atlas* Muds Raw

WDLPN1-4 106

F-1
•

••• • •

•

SECOND STAGE

ENGINE

SPIN TABLE

PAYLOAD

THUM STAGE—X-241 (Allegassy Balite* Lab.)

Thrust of altitude

Specific impulse (vac)

Total impulse •

Morning Time

Propellant

3130 pounds

250 seconds

116,400 lbs/sitc

373 mends

5.1W

FAIRING

FAIRING

PAYLOAD

EXPLOSIVE

BOLTS

SPIN ROCKETS

SPIN TABLE

The TRANSIT Program consists of four flights of satellites designed to provide extremely accurate, worldwide, all weather navigational information for use

by aircraft, surface, and subsurface vessels. The four

'vehicles are chlaillaated TRANSIT IA, IS, IIA and Mt

TRANSIT IA (Figures 1 and 2) was launched on 17

September 1959 with the remaining flight tests

scheduled in numerical order for February, April and

September 1960. The primary objectives of the program are: (a) To provide accurate reference information for POLARIS mini'e navigation accuracy;

(b)ared:. determination of satellite position through

use of payload transmitted radio signals (doppler

shift measurement); to investigate the refractive

effect of the ionosphere on radio transmissions,

and Id) acquire additional geodetic and geoMILLSTONE

BEACON

THIRD STAGE

ENGINE

GUIDANCE COMPARTMENT

RETRO SYSTEM

SPIN

ROCKETS

SECOND STAGE — AJ 10.42 (Aorolat.Ofersaral)

s

4

Thrust at altitude 7700 pounds

Specific impalas (vac) 271 wends

Total impulse (min) 870,000 lbs/sec

Burning lima 115 seconds

Propellant Liquid

curly

MIST STAGS—THOR 1118M

Thrust (LL) 151,500 pounds

Specific impulse (1.1.) 248 seconds

Specific impulse (vac) 287 seconds

Owning time 158 seconds

Propellant liquid

Moore 2. TRANSIT Me three acme PO, _vehicle

dada view of MAW NNW to ;drat stage,

Ard exPleded now of wand said third stage: mod

Paiwa •

physical information by precision frocking of the

satellite in orbit. MIND responsibilities in this program includes providing the launch vehicle, integration

of payload to launch vehicle, communications to the

ARPA tracking and data handling agencies during the

launch, and flight operations from launch through

attainment of orbit Design, fabrication and testing of

the payload will be accomplished by the Applied

Physics Laboratory for ARPA. ARPA has retained

responsibility for operating, tracking, and recording

and processing all satellite data.

F-2 WDLPM-4 106
••

t

Ilk' •

LAUNCHING SUPPORT TUBE

MECHANICAL TIMER

INSULATION

RADIATION SHIELD

ANTENNA

ANTENNA COUPLING

NETWORK

COMMAND

RECEIVER

SOLAR CELL

BANK

TELEMETER

NICKEL-CADMIUM

BATTERIES

NICKEL-CADMIUM

BATTERIES FOR SOLAR

POWER SYSTEM

NYLON LACING

OUTER AND INNS

LACING RING

SILVER-ZINC

(LIT-1001 TITAN Minn BATTERIES

INFRARED SCANNER

OSCILLATORI2)

SILVER-ZINC

BATTERIES

SOLAR CELL

BANK

DE-SPIN WEIGHT

1-SPIN CABLE

DE-SPIN WEIGHT

T LIMITER RELEASE CABLE

HANDLING RING

Pion S. Csol-aowy eosin of TRANSIT IA Payload (NAV I).

PLIGHT YOWL. DISCRIPTION

The first flight vehicle (Figures 1 and 2) consisted of

three stages, a THOR booster, AJ10-42 (Aerojet-General) liquid propelled second stage with Bell Telephone

Laboratories radio-inertial guidance systems and a

solid propellant third stage (Allegany Ballistic Laboratory X2411). Subsequent flights will use a two-stage

vehicle consisting of the THOR booster and the ABLESTAR (AJ10-104) second stage.

PLIGHT DESCRIPTION

All four vehicles will be launched from the Atlantic

Missile Range. The remaining TRANSIT payloads will

be injected into a circular orbit with a nominal altitude

of 500 nautical miles. The angle between the satellite's

orbital plane and the earth's equatorial plane will be

approximately 50 degrees on TRANSIT IB and 67.5

degrees for TRANSIT RA and IIB.

PAYLOAD ORBITAL PIMPORMANCE

The spherical TRANSIT payloads ore approximately

36 inches in diameter and weigh between 200 and 270.

pounds. Payload equipment includes four transmitters

(on frequencies of 54, 1011, 162 and 216 megacycles),

two receivers, and a gate which permits the insertion

of data only when the gate has been opened at a

previously scheduled time. Power for the first five

months will be supplied by batteries, recharged by

solar cells located in a 12-inch band around the

sphere. Extremely accurate position data can be

obtained by measuring the doppler shift of satellite

transmissions during a pass over a receiver (aircraft

or ship). Since the exact satellite position is known at

all times during the orbit, measurement of the doppler

shift provides a precise rate of time change between

the satellite and the receiver. Studies have shown that

the effects of ionosphere refraction on doppler shift

measurements can be eliminated by using the transmissions of two satellites. Navigational fixes of 0.1 mile

accuracy are expected to be obtained. During the first

three months of flight the four transmitters will be operated to obtain experimental confirmation of the *ea-

•

WDLPM-4 106 CO

►

FIDENTIAL' F-3
Table I. TRANSIT IA Powered Trajectory Flight Plan.

Vertical lift-off

Apply Stage I pitch program

Constant angle of attack

Constant attitude

Twelve sec of vernier operation

required to allow the BTL guidance

system to accumulate trajectory data

for second-stage guidance

Stage II separation plus ignition

After allowing transients to damp out

(approximately 10 sec), initiate the

BTL guidance--apply pitch rate

Apply Stage II pitch program to place

vehicle in proper attitude for orbital

injection

Constant attitude and spin-up third

stage and payload (2-3 rps)

Coast to injection altitude at constant

attitude

Constant attitude

Separate payload from Stage III

De-spin payload (2-3 rpm)

0 to 10 sec

10 to 12 sec

12 to 130 sec

130 secto Stage I main engine

cutoff (152 sec approx)

152 to 170 sec

170 sec

180 to Stage II cutoff

minus 13 sec

Stage II cutoff minus 13 sec

to Stage II cutoff minus 3 sec

Stage II cutoff minus 3 sec to

Stage II cutoff

Stage II cutoff to Stage III

ignition

Stage III ignition to burnout

After Stage III burnout

In orbit

F-4

retical mathematical relationship between the frequency and the refractive index of the ionosphere.

After four months of tracking the satellite by measurement of the doppler shift the exact position of the

satellite should be known at all times during an

orbital cycle. Once position of the satellite has been

established by precise mathematical computation,

reliable and accurate navigational fixes can be

obtained. In addition, precise measurements of the

earth's geoid can be made.

•

GROUND SUPPORT STATIONS

Tracking stations will be operated by the Applied

Physics Laboratory at Howard County, Maryland;

Las Cruces, New Mexico (University of New Mexico);

Austin, Texas (University of Texasl; Seattle, Washington (University of Washingtoni; and Newfoundland.

Second stage guidance and first and second stage

tracking andtelemetrywillbeprovidedbytteAtionticAlaulleRange.

WDLPM4 106
TRANSIT PROGRAM—MONTHLY PROGRESS

• •

TRANSIT 11-B. Aerojet-General Corporation, Azusa

Division, has been selected as the second stage

contractor. A meeting was held on 19 November at

which ground rules were established for program

management and contractual responsibilities were

defined. The first tschnkal direction meeting will be

held during December. The initial major milestone

schedule, including work on the second stage, was

established and given to the contractors. Complete

schedules will be established and distributed during

December. The delivery of required drawings and

specifications was begun during November.

Technical Progress

TRANSIT I-B. Validation of the TRANSIT I-8 checkout

van was completed at Space Technology Laboratories in Los Angeles. Subsequently, several system

and subsystem compatibility tests were completed

successfully. These tests will continue into December.

Functional tests on the nose fairing were completed

with excellent results.

* • •

TRANSIT II-A. The payload will be comprised basically of Iwo spheres whose size precludes the use

of the TRANSIT 1-11 nose faking. Studies indicate

that the COURIER I-A nose faking design can be

used on this vehicle. Additional fairings are being

procured.

• •

Delay in the release by ARM of funds for the down

range tracking and telemetry van contributed to the

30-day slippage in the launch date. However, these

funds have been released and approval was obtained

during November for design and fabrication work.

Design of all components is now in progress. Systems

tests are scheduled to be initiated during February.

Program Administration

TRANSIT I-A. The final report on the launch of this

vehicle has been prepared and will be distributed

during December. STL Report TR-59-0000-00893,

which delineates specific causes of problems encountered on this flight, also has been published. This

report also recommends certain design changes and

modifications for the TIROS vehicle as a result of

data obtained from the TRANSIT I-A.

• •

TRANSIT I-B. This navigation satellite will be placed

into a 500 nautical mile circular orbit with an inclination angle of 49.9 degrees. The launch date for this

vehicle has been rescheduled to 15 March 1960

because of second stage propulsion difficulties (see

ABLE-STAR Vehicle Monthly Progress section), and

down rang* tracking and telemetry van availability

keg technical progress section). The Detailed Test

Objective document was issued during November.

The First and Second Stage Flight Termination System document has been transmitted to the Atlantic

Missile Range for approval of the rang. safety

system. The range safety trajectory data will be completed during December.

• •

TRANSIT II-A. Launch is scheduled for 15 April 1960.

This navigation satellite payload will be placed into

a 500 nautical mile circular orbit with an inclination

angle of 67.5 degrees. A slight "dog-leg" will be

flown by the first stage, due to range safety con-

- siderations.

WDLPM-4 108 F-S
he TIROS Program consists of one flight from

he Atlantic Missile Range in January 1960. Prinary objecitves include: (a) To determine the

Feasibility of using an earth satellite to measure,

record, and transmit synoptic weather conditions; (b) To establish system parameters for

weather satellites; (c) To acquire information on

electromagnetic propagation through the atmosphere and acquisition of additional geodetic

and geophysical data by tracking a satellite in

a precise orbit. The. National Aeronautic and

Space Administration is the primary program

agency. AFBMD is responsible for supplying the

launch vehicle, integrating the. payload to the

launch vehicle, and providing communications

to the tracking and data-handling agencieifrom

launch through attainment of orbit. Payload

design, fabrication and testing will be accomplished by the Radio Corporation of America for

NASA. NASA retains cognizance for operating,

tracking, and recording and processing of satellite data.

VEHICLE DESCRIPTION

The three-stage TIROS vehicle (Figure 1) consists

of a THOR Booster, Aerojet-General (Ail 0-42/

liquid propellant second stage with Bell Telephone Laboratories radio-inertial guidance

system, and the Allegany Ballistics Laboratory

solid propellant third stage (248). Design specifications for each of the three stages are

shown on Figure 1.

,T Specific impulse (vas) 250 seconds

Total 'maybe 116.400 lbs/sec

Bursting Time 37.5 seconds

Propellant Solid

MOND STAGE •— All 0.42 (Areroist.delserial)

Thrust of altitude 7700 pounds

Specific impulse (vac) 271 seconds

Total impulse (min) 870,000 ilts/sec

urning time 115 moods

Propellant Liquid

PIOT STAGE—THOR IltilAt

Thrust (s.l.) 151,500 pounds

Specific Impulse (s.l.) 248 seconds

Specific impulse (vac) 287 seconds

urning tine. 138 seconds

Propellent

FLIGHT DESCRIPTION

The sequence of events for the powered flight

from launch at AMR to injection into orbit

given in Table 1. The payload will be placed in

a 400 nautical mile circular orbit having an

inclination angle of 51 degrees. Orbital life is

expected to be five months.

PAYLOAD OBJECTIVES

The 270 pound, cylindrical payload will be 42

inches in diameter and 17 inches in height. Payload equipment includes 2 television cameras

and an infra-red scanner designed to observe,

record and transmit weather data. Power sources

include sixty 20-volt nickel-cadmium chemical

batteries and 7344 solar cells to recharge the

batteries. The solar cells, installed in the top and

cylindrical side walls of the satellite will furnish

on average output of 13 watts for the first 140

days of vehicle life. Once during each orbit the

satellite will be interrogated and reprogrammed

from a ground station. The two television cameras have different resolution capabilities and

coverage patterns to permit observation of a

wide variety of cloud patterns. Two modes of TV

system operation are possible. When the satellite is within radio communications range of a

ground station, pictures may be taken on command and transmitted directly to earth. When

the satellite is beyond radio communication

range, camera operation is controlled by a clock

and programming circuits and the images

recorded on magnetic tape for readout during

the next pass over a ground station. The infrared scanner unit will have eleven sensors of

differing spectral responses and orientations,

permitting a comprehensive accumulation of

reflection and radiation characteristics data. The

infrared detectors operate continuously during

each orbit and the readin data is recorded on

magnetic tape for accelerated playback during

each pass over a ground station. Four beacon

transmitters are installed on the bottom side of

satellite to facilitate tracking.

GROUND SUPPORT STATIONS

The Air Force ground station at Kaena Point will

be used to support this program. Tracking and

data acquisition will be conducted on 108 ma

and command transmission on 140 mcs. Required

modifications to the TLM and VERLORT radars

TIVID STAGS—x-14$ Oil . legally Ballistic WC)

Thrust at altitude 3150 pounds

•

WDLPM-4 106 G-1
are in progress. Use of this support station will

result in: (a) minimum cost by maximum use of

existing facilities, (b) minimum equipment modification and operation effort, and (c) a satisfbctory system configuration with minimum complexity. Use of this facility also will benefit •the

SAMOS and MIDAS programs by attaining on

early buildup of experienced personnel.

TABLE I. TIROS POWERED TRAJECTORY FLIGHT PLAN

Vertical lift-off

Stage I pitch program begins

BTL guidance begins Stage I closed-loop

steering

End Stage I pitch program; end Stage I

steering.

Stage I constant attitude flight

NECO - Stage I main-engine cutoff

3.8 sec of vernier operation prior to

Stage I - II separation

Stage II separation and ignition; begin

Stage II pitch program

BTL guidance applies yaw rate command

to yaw right

BTL guidance begins Stage II closed-loop

steering

Jettison nose fairing 182.8 sec.

BTL discrete ends Stage II pitch program SECO - 10 sec.

BTL guidance ends Stage II closed-loop

steering SECO - 7 sec.

BTL guidance applies yaw rate command SECO - 6 sec. to

to yaw left SECO - 4 sec.

0-10 sec.

10 sec.

90 sec.

130 sec.

130-159 sec.

159 sec approximately.

159-162.8 sec.

162.8 sec.

166-170 sec.

172.8 sec.

BTL discrete spins up third stage and

payload - 120 rpm

Stage II cutoff - SECO; begin coast period

Stage II - III separation

Fire Stage III rocket

19. Separate payload from Stage III

SECO - 2 sec.

SECO (Approximately 273 sec.)

SECO 1.5 sec.

SECO 550* sec to

SECO 587 sec.

SECO 597 sec.

20. De-spin payload - 12 rpm

G-2
TIROS PROGRAM MONTHLY PROGRESS

Program Administration

The TIROS launch dote has been rescheduled for

January 1960 due to payload availability. Launch,

trajectory and orbit details were established during

November. Launch will take place from Atlantic Missile Range launch stand 17A on 27 January at

approximately 0625 EST. To achieve the proposed

orientation of the satellite in relation to the sun and

the earth, the launch must take place within 20 minutes of the nominal time. The vehicle will be launched

at an azimuth of approximately 90 degrees, and roll

programmed into a nominal flight azimuth of 44.5

degrees. During flight a pitch program will place the

satellite into the proper injection attitude. The orbital

path of the satellite about the earth will be nearly

circular at approximately 3110 nautical miles. The

orbital period will be 100 minutes.

• •

With the exception of the redesign of two systems

(reported in technical progress section) all system

design and development testing has been completed

and all major subsystems are undergoing pre-launch

checkout at the Atlantic Missile Range. Redesign of

the two systems is scheduled for completion with no

revision to the launch date. The final trajectory has

been completed and given to the Bell Telephone

Laboratories for insertion into the guidance system.

The final revised Detailed Test Objectives document

has been completed and distributed. A new weight

budget was established in accordance with the new

trajectory and the new BTL programming.

Tochnkal States

Based on information obtained from the flight of

TRANSIT I-A, the second stage retro-rocket system

is being redesigned to insure proper operation.

Results of a study made by Douglas Aircraft Company indicate that the use of nitrogen, and the

orientation of the nozzles in pitch and yaw, provide

adequate safety in case of a single bottle failure.

Design revisions were finalized during November

and fabrication and testing is scheduled to be completed in accordance with the present launch date.

• •

Altitude firing tests at Arnold Engineering Development Center indicate a need to revise the present

stage three/payload separation method. One solution being considered is the delay of separation

until the residual thrust of the third stage is below

the energy limits of the separation spring. Analytical

studies show that the separation can be delayed

approximately 30 minutes without seriously affecting

the attitude restraints on the payload at the lime of

separation. To give the desired time delay, and to

increase assurance of a satisfactory separation, the

Douglas Aircraft Company has been directed to

make the following changes in the stage three/

payload separation mechanism:

Provide a timer that will delay separation for

a nominal period of 25 minutes, with a tolerance of 10 percent. The timer should be started

by the same signal that starts the 'present 70

second timer.

Increase the separation spring force so that

the nominal separation velocity is approximately 4 feet per second.

Douglas also has been directed to study the following areas affecting the reliability of stage three/

payload separation:

Use of a separation command for the payload.

Effects of temperature environment on batteries and timers during the delayed separation

period.

3. Amount of tip-off imparted to the payload by

the increased spring force.

111110n01116

WDLPM-4 106
The ARPA COURIER Program consists of two flights

from the Atlantic Missile Range in May and July 1960.

The program objective is to test delayed repeater communications between a satellite and ground stations.

The program also will be used to determine the operating characteristics and capabilities of the ABLE-STAR

(AJ10.104) second stage vehicle. The program is being

conducted under ARPA Order No. 98, dated 1 July

1959 (Project Code No. 2200). AFBMD responsibility

includes development, of the launch vehide, payload

integration, launch, injection of payload into orbit,

and verification of orbital parameters at injection. The

Army Signal Research and Development Laboratory

will design, develop and fabricate the payload, and

will be responsible for world-wide ground station

requirements. Primary payload contractor is Philco

Corporation.

Vehicle Description—The two-stage COURIER vehicle

consists of a THOR booster, an ABLE-STAR (AJ10-104)

second stage and a 500 pound COURIER payload.

Booster flight control is exercised by a gyro platform

and a programmer. The second stage is controlled by

a gyro used to govern engine gimballing during powered flight. Stability during second stage coast is

provided by the "on-off" operation of jet nozzles

operating from a dry nitrogen supply. The second

stage propellants are inhibited red fuming nitirc acid

and unsymmetrical dimethyl hydrazine. The engine

will have a restart capability. The 500 pound COURIER

payload is a 60-inch sphere, containing radio repeaters, storage and memory equipment, and a battery

power source.

Flight Description—Both vehicles are to be launched

from the Atlantic Missile Range. After first stage burnout, the ABLE-STAR vehicle will place the payload into

the desired trajectory and then shut down. The second

stage and payload will coast to the desired 650 nautical mile orbital altitude and the ABLE-STAR engine

reignited to attain orbital velocity. The orbital angle

of inclination will be 28.5 degrees from the equatorial

plane. The orbital period will be 110 minutes.

Payload Objectives—Storage and memory elements

in the payload will deliver messages, upon command,

to each of three ground stations; as well as exchanging "real time" information when the satellite is within

line-of-sight of two ground stations. During these periods a ground station can relay messages direct to the

next ground station, through the satellite simplex

repeater equipment.

Ground Support Stations—These stations will be

located at Camp Salinas, Puerto Rico; Torrejon Mr

Force Base, Madrid, Spain; and Holman°. Hawaii.

Station design and development is under contract to

International Telephone and Telegraph Corporation.

WDLPM-4 106

UCOND STA0T—AILT4TAR (A)10.104)

Thrust at altitude

Specific impulse (voc)

Tafel kopeks (min)

Morning lime

Propellant

$030 pounds

278 seconds

2.3 X 10' lbasec

294 se ands

Liquid

NUT STA01--T1101 SUM

Thrust 0.1.1

Specific impulse (1.14

Specific impulse (vac)

Burping Hoe

Propellant

151.500 pounds

24$ woods

267 same

138 seconds

Liquid
-CONFIDENTMl

COMM PROGRAM—MONTHLY PROGIRMIS

Program Administration

COURIER I-A. The launch date is 15 May 1960. The

payload will be placed into a 650 nautical mile

circular orbit with an inclination angle of 28.5

degrees.

• •

COURIER I-B. Aerojet-General Corporation, Azusa

Division, has been selected as the second stage

contractor. A meeting was held on 19 November, at

which ground rules were established for program

management and contractual responsibilities were

defined. The first technical direction meeting will be

held during December. An initial major milestone

schedule, including work on the second stage, was

established and given to the contractors. Complete

schedules will be established and distributed during

December. The delivery of required drawings and

specification: was begun during November.

Technical Progress

COURIER I-A. Validation of the checkout van was

completed at Space Technology Laboratories in Los

Angeles. Subsequently, several system and subsystem

compatibility tests were completed successfully. These

tests will continue into December.

• •

Following the release by ARPA of the necessary

funds, work was initiated on the down rang. tracking and telemetry van. Design and fabrication

approval was obtained and the design effort has

been started. Systems tests are scheduled to begin

during February.

WDLIPM-4 106 H4
A G E NA

Figaro 1.

Drawls; of AGENA "B..

propoltiosr system

• • .1. •••

The AGENA vehicle was first designed by the Air

Force as the basic satellite vehicle for the Advanced

Military Reconnaissance Satellite Program (Weapon

System 11711 The vehicle was designed to be

boosted by an ATLAS 'ICBM, and basic AGENA

dimensions were derived from this booster selection.

Abe type of trajectory possible with the ATLAS

booster, coupled with the stringent eccentricity

requirements of the program, led to satellite guidance selection suited to orbital injection in a horizontal attitude. This in turn led to the development

for the AGENA of the horizon scanner attitude control system. ffhe Bell Aircraft L181-Ba-3 engine

(Bell-Hustler rocket engine developed for the 6-58

Bomber) was chosen for AGENA propulsion sines it

was in an advanced state of development. Subsequent modifications to this engine to change the fuel

from JP-4 to .UDMH for increased performance

resulted in development of the XLR-81-Ba-5 engine.

In an effort to accelerate flight testing of the AGENA

stage and subsystems, the low cost and early availability of the THOR IRBM caused a study of this missile as a booster for the AGENA to be made. When

results indicated that the AGENA could be successfully flight tested in low altitude orbits, utilizing the

THOR as the first stage, the DISCOVERER Program

was initiated. This program has resulted in five

WDLPM-4 106

launches of the AGENA vehicle to date, three of

which attained orbit. The last two launches resulted

in very satisfactory performance of the entire vehicle,

including the AGENA stage and all subsystems. The

only malfunction noted was failure of a battery

designed to activate the recovery aids of the recoverable capsule fitted to the vehicles for biomedical

research purposes. Satellite propulsion, guidance,

stabilization, and auxiliary power systems functioned

as programmed. As a result the AGENA vehicle has

become a flight-tested and proven vehicle in advance

of the date it will be needed to place advanced mil- ,

itary reconnaissance payloads into operational orbit

with the use of ATLAS boosters.

The AGENA vehicle forward section contains satellite

guidance, power conversion, telemetry, and command equipment. The aft equipment section supports

the engine thrust vector hydraulic control system and

the pressure storage vessels for propellant pressurization and attitude controls. The center section consists of propellant tankage. The vehicle attaches to

the booster by means of an adapter section surrounding the oft section.

Progress in the design of payloads for the MIDAS

Program developed an urgent need for higher altitude orbits. This resulted in a program for development of the XLR81-Ba-7 engine, which has a single

1-1
NOSE CONE

FUEL AND

OXIDIZER

TANK

SECTION

SEPARATION

PLANE

ELIUM

SPHERES

NITROGEN

SPHERE

ULLAGE ROCKETS

NGINE (UDMH)

NOSE CONE

FUEL AND

OXIDIZER

TANK

SECTION

SEPARATION

PLANE

ELIUM

SPHERES

ULLAGE ROCKETS

NITROGEN

SPHERE NGINE (UDMH)

restart and extended burn capability. This capability

permits placing a given payload weight on a higher

orbit. This development eased the payload restrictions in use of the extremely reliable THOR as a

booster for AGENA. The AGENA was redesigned to

double the copeck, of the propellant tanks, providing increased performance.

Payloads may be installed either as a package on

the forward equipment rack or distributed throughout the AGENA vehicle. Payload capability is as

much dependent upon available space, within the

vehicle as upon propulsion energy available. Generally, payloads in excess of 1,500 pounds must be

integrated throughout the AGENA stage. A program

for development of maximum efficiency of the first

stages was also underway. The ATLAS ICBM was

modified to reduce limitations imposed by hetedory

requirements and static load capability.

AGENA "A" AGENA "a"

Weight—Wet 1,400 Weight—Wet 1,450

Fuel (UDMH) 1,900 Fuel (UDMH) 3,800

Oxidizer (IRFNA) 4,800 Oxidizer (IRFNA) 9,600

GROSS WEIGHT 8,100 lbs. GROSS WEIGHT 14,850 lbs.

;.-

J-2 NNWCONIFID ilIDLYM-4 106
"•44A1P9r-s. *r , r.

thlt•Cal...if

•ti..,

2 • ,

AGENA VEHICLE—MONTHLY PROGRESS

Centrifugal testing of the AGENA II" test article

propellant tanks has been completed satisfactorily

and vibration testing is now in progress. Tank design

stressed use of components proven in the AGENA

"A". For manufacturing simplicity, hemispherical

ends were made uniform except for material thickness. This permitted the use of straight cylindrical

sections between end pieces, which are machined

fiat, then rolled into cylinders.

• •

The AGENA "A" engine (8084) will be used for

test procedures until the AGENA "8" engine (8081)

becomes available. The 8081 will incorporate two

solid propellant turbine starters and necessary electrical modifications. The main pressure regulator fuel

and oxidizer vent devices have been redesigned,

developed and produced for the propulsion system.

This unit will be tested during the propulsion test

vehicle assembly program now in progress at the

Lockheed Sunnyvale facility.

• •

Massachussetts Institute of Technology progress on

development of an advanced orbital attitude control system for the MIDAS/SAMOS and AGENA

configuration was reviewed in November. Pitch control appears to pose no immediate problem, however, roll-yaw, controlled mainly by a large gyro,

will present difficulties for the following reasons:

The gyro size depends to some extent on the

vehide inertias which are not constant at this

time.

A survey indicates that even the most suitable

gyro available will require considerable modification.

Kearfoft has been selected as the gyro contractor.

Pigare 3. Shale gage testis: of AGM! "B" Mesa&

asperity) /set

Now 4. Phial system check.

oat Inman, heists coxlicted

ox slot fixel AGENA

vehicle (1038). All sakefsums

vehicles will be AGENA 'B"

coofigwratiox.

WDLPM-4 106
MIES

„taw..

Able Star

This program will develop a versatile and efficient upper stage for use with varied booster/vehicle

combinations. This stage will have basic design features proven in the dependable Ail 0-101 stage

used on the THOR/ABLE vehicles. Improvements

being made include: (a) increased propellant capacity; (b) a multiple restart capability; and (c) a fulltime attitude control system to operate during coast

periods as well as powered flight. These improvements will permit a two-stage THOVABLE-STAR vehicle to attain weight/altitude performance equal to

that of the three-stage THOR/ABLE vehicles. This

will provide increased reliability and accuracy.

This stage will be suitable for mating to THOR,

ATLAS or TITAN space boosters and can be

modified to accept a solid propellant third stage,

if needed.

12igsre 1. Side view of ABLE-STAR vebide wawa is &wan dolly.

WDLPM-4 106 —69/"ENTIAL— K-1
ARLE-STAR VEHICLE--MONTHLY PROGRESS

A requirement for five test runs of 300 seconds duration was established to provide a demonstration of

full duration propulsion system performance.

Although the ABLE-STAR (AJ10-1 04) thrust chamber

is practically identical with the earlier AJ10-40 and

All 042 series engines, this additional requirement

is necessary since the earlier models were tested for

only 115 seconds duration. These full duration runs

must be performed successfully prior to deliver/ of

the first Right unit.

• •

The full duration test runs were initiated during

November. Injector cooing problems, caused by

injector manifold design and use of IRFNA as a coolant, resulted in burn-through of the injector plate

and the cooling tubes dose to the injector. These

problems subsequently were solved and two thrust

chamber assemblies were operated successfully for

full duration, 300 seconds periods. Three additional

thrust chamber assemblies are scheduled for full

duration operational tests.

• •

The Preliminary Flight Rating Test engine was

accepted by the Air Force on 19 November. The first

Right test engine also passed acceptance testing.

However, the propellant tanks on this unit were subsequently damaged beyond repair due to the contractor's use of improper decontamination procedures.

• •

As a result of the problems reported in the preceding

two paragraphs the delivery schedule of ABLE-STAR

units has been revised as follows:

first Fight Unit 31 December

Second Right Unit 1 February

3. Third Right Unit 1 March

The Preliminary Right Rating Test Program is scheduled for completion by 4 January. Right schedules

for vehicles using this propulsion system are given in

the TRANSIT and COURIER Program sections.

WDLPM-4 106
ATLAS BOOSTER

MANNED SATELLITE

HORIZON SCANNER

PITCH & YAW JETS

•

ESCAPE ROCKET

POSIGRADE

RETROGRADE

JETTISON

ROCKET

RETRO & POSIGRADE ROCKETS

ROLL JETS HEAT SHIELD

ESCAPE ROCKET ENTRANCE HATCH

PARACHUTE

WEIGHT AT SEPARATION APPROX. 2413 LBS. ORBIT INCLINATION 33 DEGREES ORBITAL ALTITUDE 105-120 MILES (n) HEAT SHIELD ABLATIVE OR SINK ORBITAL CYCLES 3-18 RECOVERY AIR &/OR WATER OR LAND

Figure 1. Complete vehicle (top view) with satallits iostailmi ors ATLAS boomer. Maw" satellite

(bottom view) showing pilots' flight parities, aysd detail views of retro end posigrade rockets dad pilot safety system escape rockets.

WDLIIM-4 101Amoptogat L-1
/

/

/

/

AO BALLISTIC TRAJECTORY

TO SIMULATE REENTRY CONDITIONS

0 NEAR ORBITAL INSERTION, IMMEDIATE

REORIENTATION, DECELERATION, REENTRY

ATLANTIC RECOVERY

CI FULL ORBIT 3-18 ORBITS 105-115 N.M. WEST COAST REENTRY ATLANTIC RECOVERY

LEGEND

RESPONSIBILITY:

AFBMD

NASA

ATLAS

BOOSTER

►

I /

I /

/

/

0111116.. /

/

/

/

/

/

ESCAPE ROCKET

CAPSULE

Project MERCURY represents the transitional th. rill mg' ten ATLAS boos rs te modified in accordance;

between this nation's cumulative achievements in space

research and the beginning of actual space travel by

man. The primary program objective is to place a

manned satellite into orbit about the earth, and to

effect a controlled re-entry and successful recovery

of the man and capsule (Figure 1). Unmanned ICBM

trajectory and near-orbital flights, and unmanned

orbiting flights will be used to verify the effectiveness

and reliability of an extensive research program prior

to manned orbital flights (Figure 2). The program will

be conducted over a period of nearly two years. The

initial R&D flight test was accomplished successfully

in September 1959. The total program accomplishment is under the direction of NASA. The primary

responsibility of AFBMD to date consists of: (a) pro-

_

with program objectives and pilot safety factors, and

(b) determination of trajectories and the launching

and control of vehicles through injection into orbit.

The division of responsibilities for this program is given

in Table 1. Specific details of AFBMD support are given

in Table 2.

Major contractors participating in the AFBMD portion

of this program include: Space Technology Laboratories, systems engineering and technical direction;

Convair-Astronautics, modified ATLAS boosters;

GE/Burroughs, ATLAS guidance equipment; and Rocketdyne, engines. All of these companies also provide

special studies and engineering efforts peculiar to

meeting Project MERCURY requirements.

..4

Span 2. Plight test trajectories for PI001$ MERCURY, defining specific objectives. Trajectory C

represents the patb of the fund (semeneti) flight. The Mot at wind, APBMD and NASA responsibility

it divided represents injection into orbit.

L-2 gm& --CDROM, 'WDLP11-410e
;.• • • • • :

PROJECT MERCURY

NASA ResponsibRhy

Booster-Capsule integration

Capsule

Ground support

Recovery

Flight 1 Flights

2-10

Laboratory Flight

Capsules Capsules

(McDonnell)

Re-entry

heating,

dynamics,

recovery

AMMO RespensibMty

Booster

ATLAS RD"

Modified

New guidance antenna.

Thin skinned tanks

Shifted tank bulkheads

Light weight thrust structure

Capsule separation mechanism

Interface electrical connections

LOX dome protection

Programmer Modified

Range safety destruct time delay

Abort-sensing system

Detects dangerous variations

Launch in tank pressures, guidance,

thrust power, attitude. Can

initiate escape sequence

Integrated countdown—including pilot insertion

and capsule countdown. AMR launch Stand 14—

adapted for pilot and capsule.

GE ground guidance equipment modified to

provide: (a) real time data, (b) additional discrete

command signals during powered flight.

Launch azimuth:

Ballistic trajectory-108 •

Orbital trafectory---072•

Unmanned non-orbital flights

Manned and unmanned orbital flights

Launch from AMR

Earth track 33 •N/33 •S/Lat. (for orbital

flights)

3-18 orbits

Orbit capability 28 hours

Orbit weight 2500 lbs.

Orbit altitude 105-115 n.m.

Automatic pilot safety equipment with

manual backup

Table 1. Ow* of NASA aid AP13MD raPtiatibilisies is PROJECT MERCURY.

WDLPM4 106 L-3
' •

1,4

PROJECT MERCURY

AFBMD Responsibility

in support of

PROJECT MERCURY

NASA HS-36

includes:

Design, engineering studies

Equipment modification

Hardware fabrication

Flight scheduling

Launch support

Trajectory data

Missile allocation

Provide ten (101

ATLAS boosters.

Modify boosters for NASA preliminary research and manned orbital

flight and safety objectives.

Launch, control and define trajectories

of booster-capsule vehicle up to, and

induding, injection into orbit.

Toirie 2. AFRAID respotuihilisios in 'ippon of PROJECT MERCURY.

PROJECT MERCURY--A1ONTHLY PROGRESS

Program AcIminisfrallen

NASA has requested the addition of four ATLAS

boosters to the present nine flight test vehicle Project MERCURY Program. These additional launches

would extend the overall program through February

1962. Final agreement on the expanded program

is pending. Originally consisting of ten flight test

vehicles, the program was reduced to nine when

the second flight test was cancelled in September.

Mb flight was cancelled because of the high level

of success in attaining the test objectives of the first

flight (Big Joe I). The ATLAS booster for the cancelled

flight was subsequently transferred to the ABLE-4

Program. The remaining eight currently scheduled

flight test vehicles will consist of modified ATLAS

boosters and Flight Capsules (McDonnell) capable

of sustaining human life.

Technical Progress

The next Project MERCURY flight test is scheduled to

occur in June 1960. Designated MA-1, this flight

will carry a McDonnell Aircraft Corporation capsule

on a trajectory similar to that used on the Big Joe I

R&D test. The capsule will be boosted by ATLAS SOD,

modified to meet Project MERCURY Program requirements. On board will be components of the Abort

Sensing System which is under development to insure

pilot safety on manned flights. These components

will be tested open loop. Closed loop tests of the

complete Abort Sensing System will begin with the

third ATLAS/MERCURY flight test, designated MA-2

and scheduled for August 1960. Components of the

Abort Sensing System also will be tested on S ATLAS

R&D flights to provide additional data prior to use

on a marultd flight vehicle.

WDLPM-4 106
724'sviezo

mate CAVE CONTANEK

Af0S6E AfarThr

72 . SN/11.0

S4461/PEAC014 U6NT4 SOFAI

ISOAIIIS,E7C.

Af066E

SWAGE

ANIE1/114 hf/I/N6

4490101.4?

APICOPIA7H0d4 JPLACES

/11 CINIIE COM/NIA'

f C "ND ANTENNAE-6 PLACES

ANTENNAE -6 PLACES

P/SCONIECT ACCESS HATCH

ANTENNAE, wiltemarram

V 710WOCOOPII DISCONNECT

AfTERBCPY

CAPSLCE SEAL/NC POW

/1/STITIAIENTATION

UNA/I/CAL REaPDICIE

HATCH

;.- -

MIL/CAL efeZMACLE

moat

ADAPTER SEA\

APAPTERa...1

ABLATION NES

PIESSO/If CAPSULE

Aff/R/DE 3ETS

NITAPOCEN AMAION/A

RUM ACCESS

74,50 MAX 0.P ASLAIION SN/ELP —

N/TIVSEN 'arras

Pier* 3. Drawing "Big Joe capsal• specossfally

tuts/ os first 'MERCURY' light tat.

WDLPM-4 106 L-5
TutiFtEubtThe rough draft of the Detailed Test Objectives Document for ATLAS SOD is being coordinated at the

present time. Formai publication is anticipated by

February 1960. Current test objectives for flight

MA-1 are as follows:

Evaluate the performance of the capsule and

its systems on a simulated re-entry from orbit.

Determine the capsule full-scale motions during

a simulated re-entry from orbit.

Determine the performance of the ablation

shield and measure the afterbody heating of

the capsule during simulated re-entry from

orbit.

Establish the adequacy of the capsule track

and recovery procedures.

Establish the adequacy of the capsule landing

systems.

Recover the capsule.

Evaluate the loads on the capsule during the

actual flight environment.

Establish prelaunch checkout procedures for

the flight capsule.

• • •

On 13 November at Langley Air Force Bass, Virginia,

a joint AFBMD/Space Technology Laboratories team

presented a briefing for NASA on the AFBMD proposed Booster Quality Assurance and Safety Program. In addition to informing NASA of AFBMD

actions related to booster responsibility, the briefing

suggested that an overall capsule/booster safetyof-flight inspection be initiated by NASA. AFBMD

safety actions were endoned, and NASA indicated

that a conference would be convened early in

December to consider action on a safety-of-flight

inspection for the complete Project MERCURY system. Convair-Astronautics has been directed by

AFBMD to begin implementation of all aspects of

this program related to booster responsibility.

WDLPM4 106
iRSi

20'

30'

,.-ABL X248 A5

H202 JETS

X254 Al --N.

H2 02 JETS

0' 30"

PROJECT 609A

Spore Environment Test System

PROGRAM DISCRIPTION—The Hyper-Environment

Research Program (609A) is divided into R&D and

Operational phases. The R&D phase will be used to

develop and flight test vehicles capable of carrying 50

to 1,020 pound payloads to altitudes of 200 to 7,000

miles. The Operational phase will use this standardized

vehicle to permit the economical performance of flight

test experiments in support of scientific research and

advanced military space system programs.

Rconomy—Rollabillty—Versatility—In this order

of emphasis ore the three significant guides to program accomplishment. ECONOMY is being achieved

: • ,

THIOKOL

SERGEANT

(IMPROVED)

XM33E5

(ALT. NOZZLE)

H2 02 JETS

31"

AEROJET

SENIOR

33K S 120,000

JET VANES AND

AERO FINS

ABL X254 Al

THIOKOL

SERGEANT

XM33E3

S.L. NOZZLE

JET VANES AND

AERO FINS

Pigmy I. Three vorietioes of Project 6094 vehicle

deorosstratethe otistiots-versetility of the program

VITDLPM-4 /06 M-1
Pigare 2.

Pow difersei

ersisetoria

oossihk

difantsinamar

of Pro** 6094

amps.

M-2

by long range planning and maximum integration with

other programs. Use of the bask four-stage, solid propellant, SCOUT vehicle, developed by NASA and

modified to achieve Program 609A objectives, will

effect an economy in vehicle development. Necessary

modifications include provisions for stabilizing the

fourth stage without spin and use of the vehicle in less

than the full four-stag. configuration. Close integration with the current ballistic missile program will effect

an economy by permitting tests and experiments to be

conducted on regularly scheduled ballistic missile test

flights whenever possible without delaying schedules.

Economy in the operational phase will be exercised by

the use of this low-cost vehicle as a standard flight test

platform to perform scientific and military experimental research in support of all Air Force facilities. REUABIUTY will be obtained by a nine or ten vehicle R&D

flight test program, at least four flights of the bask

SCOUT, and maximum use of knowledge gained in

prior Air Force ballistic missile flight testing. VERSATIUTY will be achieved by designing a vehicle capable

of being readily adapted to a wide range of payload

variations, and capable of being flown in several configurations of four stages or less. This VERSATILITY

results in the following flight capabilities: (a) vertical

probes having a wide variance of payload weight/

attitude combinations; (b) boost-glide trajectories; (c)

ballistic missile trajectories; (d) downward boosted,

high-speed re-entry profiles, and (e) full orbit to

approximate maximum of 400 miles with 150 pound

payloads.

Program Management —An abbreviated development plan, covering the R&D phase only, was

approved on 9 January 1959. Funds in the amount of

68,180,000 were made available for this abbreviated

portion of the program only. A letter was issued assigning management responsibility to ARM, with emphasis on integrating the program with the scientific and

military research experiments conducted on regularly

scheduled ballistic missile flight tests (Piggyback Program). In June 1959, Aeronutronic Division of the Ford

Motor Company was chosen through normal competitive bidding as the Payload, Test, and Systems integration Contractor. Arrangements have been made for

the procurement of vehicle components and associated

support equipment, modified to meet Program 609A

requirements, through NASA, rather than through the

SCOUT Program contractors. Atlantic Missile Range

facilities consisting of Launch Pad 18A, its tower, and

half of the blockhouse will be made available to the

Air Force for this program. A division has been established within the 6555th Guided Missiles Group at

AMR to supply Air Force technicians to participate in

the assembly, checkout and launch operations of the

R&D phase under the direction of the Payload and

Test Contractor. An all-military operational capability

will be developed from within this group.
ore • M.

PROJECT 609A MONTHLY PROGRESS

Hq USAF (AFDAT) approval of the 609A Development Plan carried the stipulation that launch schedules be arranged to gain maximum advantage from

the results of NASA SCOUT test launches and from

experiments carried on 609A R&D flights. USAF also

stated that there is no firm date by which the 609A

development-test program must be completed. This

stipulation was significant, since the original flight

schedule would compromise both objectives. in

accordance with Hq USAF instructions, 609A program schedules have been revised to better attain

both objectives.

• •

Launch of the first NASA SCOUT vehicle is scheduled

for 15 January 1960, approximately three months

behind the originally scheduled mid-October date.

Further delays in this date as well as in the frequency

of subsequent launches may occur. This delayed

schedule has caused a similar effect upon the 609A

flight scheduling.

•

Test stand 18B at the Atlantic Missile Range will be

available for Project 609A about 15 March 1960.

This stand has air conditioning %daft*, required for

payloads during countdown, adequate pad space

for maneuverability, greater hoisting capability, and

required electrical facilities.

• •

A Design Review Conference, held at NASA Langley

Research Center on 17-19 November, was attended

by representatives of all organizations involved in

determining the design concept of the 609A family

of vehicles. AFBMD and NASA contractors have

completed design modification studies sufficiently

to permit recommendations to be made regarding

609A vehicle configurations. USAF decisions were

made regarding 609A modifications to the SCOUT

vehicle. The decisions were based on compatibility

of SCOUT design with 609A requirements, 609A

performance and growth potential, modification cost

and time required, and the desire for a cooperative

USAF-NASA effort in developing an interchangabie

family of research support whiles. NASA has

decided to incorporate many of the USAF 609A

modifications in its future SCOUT vehicles, decreasing the differences between the two vehicles.

• •

The 609A vehicles will be launched by Air Force

personnel assigned to the 6555th Guided Missile

Group, presently attached to the 609A Test Office

at Patrick Air Force Base. A familiarization program

was conducted for this group at the contractor's

facilities from 9 November to 20 November. ARDC

personnel will receive short-term, periodic familiarization tours throughout the flight test portion of the

program. The major portion of the experience and

proficiency In assembly, checkout and launch (by

military personnel) will be attained by actual participation, under Aeronutronic supervision, during the

Sight-test phase of the program at AMR.

• •

Hq ARDC (RDRRB) has requested the 609A Project

office to compile, validate and integrate the proposed experiments from the ARDC Centers for the

Phase II (operational) portion of the program during

the interim period of ARDC reorganization. This

function was formerly performed by the various RDR

Staff Offices at ARDC.

On 1 October, it was learned that the entire FY 1960

funding program has been deferred by the Office of

the Secretary of Defense. This will have no impact

on the program provided funds are released prior

to 15 December. After this time serious program

slippage will occur. This information was transmitted

to ARDC with a request for action to obtain early

release of the funds. On 19 October, a message was

received which authorized initiation of the funds.

Although authority for commitment and obligation

was specifically excluded, this action indicated that

complete release of the funds may be expected soon.

As of 30 November, FY 1960 funds remain frozen.

WDLPM-4 106 M-3
SPACE

studies
ADVANCED SYSTEMS STUDIES

The Advanced Systems Studies Division has several space studies in progress. The purpose of these

studies is to determine the military missions and

mode of operation in space. For the purpose of

study, space has been divided into three brood

areas; earth orbital, lunar, interplanetary. Studies

in the lunar and interplanetary area are being managed and directed at AFBMD. There are two studies

in the Lunar area: SR 192 (U) Strategic Lunar System, and SR 183 (U) Lunar Observatory. There is one

study in the interplanetary area: SR 182 (U) Strategic

interplanetary System.

The objective of SR 192 is to determine a military

posture in the lunar area which is defined as the

surface of the moon and the area in its surrounding

gravitational field. This is a broad conceptual type

study which will examine all facets of military operations such as offensive, defensive, and supporting

systems. This study was funded with $600,000 in

Fiscal Year 1959 and final reports from the contractors are due at AFBMD by February 1960. In

.• •

logical approach for establishing a manned intelligence observatory on the moon from which the

entire earth and its surrounding area can be kept

under continuous surveillance. All earth orbital systems can be monitored and enemy activities in space

and on the lunar surface can also be watched. All

possible types of sensors and their probable ranges

will be examined. This study will also include the

means of logistically supporting and establishing the

lunar base. This study was funded with $420,000 in

Fiscal Year 1959. Three contractors were funded

and three additional contractors are performing the

study on a voluntary basis. Consequently, it is estimated that this study has the equivalent of 51.5

millions being applied to it.

4. The interplanetary area is being studied under

SR 192 (U) Strategic Interplanetary System. The

objective of this study is to determine the possible

military missions and the type of equipment necessary for operations in the interplanetary area. This

• • •
ADVANCED SYSTEMS STUDIES

The Advanced Systems Studies Division has several space studies in progress. The purpose of these

studies is to determine the military missions and

mode of operation in space. For the purpose of

study, space has been divided into three broad

areas; earth orbital, lunar, interplanetary. Studies

in the lunar and interplanetary area are being managed and directed at AFRMD. There are two studies

in the Lunar area: SR 192 (U) Strategic Lunar System, and SR 183 (U) Lunar Observatory. There is one

study in the interplanetary area: SR 182 (U) Strategic

Interplanetary System.

The objective of SR 192 is to determine a military

posture in the lunar area which is defined as the

surface of the moon and the area in its surrounding

gravitational field. This is a broad conceptual type

study which will examine all facets of military operations such as offensive, defensive, and supporting

systems. This study was funded with $600,000 in

Fiscal Year 1959 and final reports from the contractors are due at AF8MD by February 1960. In

addition to the three funded contractors working on

this study, there are three voluntary contractors.

Consequently, the total effort being applied is estimated as equivalent to one million dollars.

3. An obvious military requirement in the lunar area

will be a surveillance and intelligence collection system. Therefore, SR 183 (U) Lunar Observatory was

initiated to examine this problem. The objective of

this study requirement is to determine a sound and

logical approach for establishing a manned intelligence observatory on the moon from which the

entire earth and its surrounding area can be kept

under continuous surveillance. AR earth orbital systems can be monitored and enemy activities in space

and on the lunar surface can also be watched. All

possible types of sensors and their probable ranges

will be examined. This study will also include the

means of logistically supporting and establishing the

lunar base. This study was funded with $420,000 in

Fiscal Year 1959. Three contractors were funded

and three additional contractors are performing the

study on a voluntary basis. Consequently, it is estimated that this study has the equivalent of $1.5

millions being applied to it.

4. The interplanetary area is being studied under

SR 192 (U) Strategic Interplanetary System. The

objective of this study is to determine the possible

military missions and the type of equipment necessary for operations in the interplanetary area. This

area is being studied separately from the lunar area

because the operational problems involved appear

to be somewhat different, the distances are much

greater; our present knowledge of the area is limited, therefore, special types of navigational and

propulsion systems will be required. This study was

funded with $285,000 in Fiscal Year 1959 which

has been distributed among three contractors. Contractors' final reports are due at AMMO in February 1960.

WDLP64.4 106 N-1
DISTRIBUTION

Headquarters, United States.Air Force 19

Headquarters, Air Research and Development Command 10

Strategic Air Command 1

Air Force Cambridge Research Center 1

Air Force Flight Test Center 1

Rome Air Development Command 1

Air Force Missile Defense Command 1

Wright Air Development Center 1

Air Force Special Weapons Center 1

Mr University 4

Arnold Engineering Development Center 1

Air Proving Ground Center 1

.Air Defense Command 1

Air Training Command 1

Air Photo. and Charting Service 1

Air Force Missile Test Center 1

United States Air Force Academy 4

Air Technical Intelligence Center 1

Assistant Commander for Missile Tests 2

Air Force Ballistic Missile Division (ARM) 60

Ballistic Missiles Center (AMC) 12

Assistant CINCSAC (SAC MIKE) 3

WDLPM-4 106