Infrared Payload Tracking Tests

@ An operational test of the complete MIDAS sys-
fem was conducted on 29 January. The
test target wos an ATLAS missile launched from Van-

ground data link system transmitted the dato to the
tracking station where it was tape recorded. Analysis
of the tapes on which the tracking information was

recorded revealed that the target information ob-
tained was highly satisfactory. The capability of the
space and ground presentation equipment, as in-
stalled at Vandenberg AFB, was established. The test
also provided a valuable Personnel training func-
tion. On 4 February, the launch and flight of the
DISCOVERER IX vehicle was tracked in a similar
manner for 110 seconds.

@ Advanced Presentation Unit — Negotiations
between LMSD and General Electric Co. on the con-
tract for this unit are essentially complete.

@ Solar Auxiliary Power — Fabrication of the
solar array panels was started on 8 February. The
mockup of the entire array is nearly complete. A
functioning 1/10 scale mode! of the array mech-
‘anism was completed during February.

© Reliability Negetiations — LMSD and Bell Air-
craft have completed a work statement for an
AGENA engine relicbility program.

Fignre 9. Infrared scanner weit and associated equipment
shows in test set-up at Vandenberg APB. Infrered bracking
of ATLAS and DISCOVERER flights were berformed with
Aighly successful results. Note the bilser-ont position of the
scanner, Satellite borne communications system cirenitry is
shown at right side of photographs.



—EGONFIDENTIAL _

The TIROS Program consists of one flight from the
Atlontic Missite Range early in 1960. Primary objec-
tives include: (a) To determine the feasibility of using
Gn 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 geo-
detic 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 vehi-
dle, and providing communications to the tracking
‘and data-handling agencies from launch through
‘attainment of orbit. Payload design, fabrication and
testing will be accomplished by the Radio Corpora-
tion of America for NASA. NASA retains cognizance
for operating, tracking, and recording and process-
Ing of satellite data.

VEHICLE DESCRIPTION

The three-stage TIROS vehicle (Figure 1) consists of
@ THOR Booster, Aerojet-General (AJ10-42} liquid
Propellant second stage with Bell Telephone Lab-
oratories. 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.

9

THIRD STAGE—X-248 (Allegany Ballistic Lab.)
{| Thrwst of oltHude 3150 pounds

Specific impulse (voc) (250 seconds

Total impulse 116,400 Ibs/sec

Burning Time 37.5 seconds

Propeltont Solid

SECOND STAGE—AJ10-42 (Acrejet-General)

Thrust of altitude 7700 pounds
Specific impulse (vac) 271 seconds
Totel impulse (min) 870,000 ths/sec
Burning time 115 seconds
Propellant liquid

FIRST STAGE—THOR IRSM

Thrust (0.1) 151,500 pounds
Specific impulue (1) 2481 seconds
Specific impulse (voc) 287 seconds
Burning tine 158 seconds
Propellant Uquid

WDLPM-4 188

ay)

. The sequence of events for the

Powered flight from launch at AMR
to Injection into orbit is 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 equip-
ment includes 2 television cam-

=

eras 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 cylindrica! side walls

e
|
|e
of the satellite will furnish on aver-

age output of 13 watts for the first

140 days of vehicle life. Once during each orbit the
satellite will be interrogated and reprogrammed from
@ ground station. The two television cameras have
different resoluiton capabilities and coverage pat-
tems to permit observation of a wide variety of cloud
patterns. Two modes of TV system operation are
possible. When the satellite is within radio commu-
nications range of a ground station, pictures may be
taken on command and transmitted directly to earth.
‘When the satellite is beyond rodio 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
ever a ground station. Four beacon transmitters are
installed on the bottom side of the satellite to facili-
tate 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 mes and com-
mand transmission on 140 mcs, Required modifica-
tions to the TLM and VERLORT radars are in progress.
Use of this support station will result in: {a) minimum
cost by maximum use of existing facilities, (b) mini-
mum equipment modification and operation effort,
and (c} a satisfactory system ‘configuration with mini-
mum complexity. Use of this facility also will benefit
the SAMOS and MIDAS Programs by aitaining an
early buildup of experienced personnel,

CONFIDENTIAL

Hl
4


(seq)

F4

Time
(sec) Stage Event Stage Event.
X+01 | Liftoff switch activates X+ 2300 it Jettison nose fairing
Programmer ‘starts Times vary in fl Stop pitch Program
Gyros uncaged accord. with Pitch command
fl Umbilicals eject trajectory. Yaw command
Arm destruct initiator X ~ 429.0 lt Stop pitch program
X+2 | Roll program initiated X— 44s It Engine cut-off signal
X+9 | Roll program complete Switch pneumatic coast control
x+10 | Pitch program initiated *yatem on pitch, yaw and roll
Ist step pitch rale As required xX- 4710 u Tun off hydraulic power
x +25 1 "Bnd step pitch rate for trajectory Xo 00 nitiate coast phase pitch program
x+#70 1 _3ed step pitch rate and detailed in X— 10368 Stop coast have pitch program
X40 1 “Wh step pitch rate DTO Sa 1686 Start hydro Power
. [Autopilot gain chonge X— 14886 Il Engine restart fire signal
tt“ Progrommer armed Uncage accelerometer
xP 10 1 Pitch program com plete X— 14916 MN Cage coast pneumatic control system
xF 182 | Main engine cut-off (MECO) X 1500 Ihe TS catolfprobe (back-up)
Circuitry armed Arm oxidizer probe (back-up) _
X+1635 1 MECQ backup ormed Arm spin and separation mechanism
- xX +180 | MECO X — 1504.0 " Engine cut-off signal
“xX 70 Stor programmer Uncage coast pneumatic control system
X +1700 (Engine fire signal Start spin table
Uncage thrust chamber Start timer on spin table
in pitch and yaw X — 1506.0 ul Engine cut-off [bock-up)
Uncage high thrust roll jets Start spin table {back-up}
Uncage gyros X — 1507.0 ||___ Remove spin table bolt power
X+ 170.85 Il Blow separation bolts X— 1526.0 It Blow separation bolts
X+ 1760 It Start pitch program Activate separation actuators
Payload separation occurs
Sequence of Events—TRANSIT 1B, 2A and 28
en er $e



‘WDLPM-4 188

MONTHLY PROGRESS—AGENA Program

@ All of the development efforts telated to the
AGENA “BY double capacity integral propellant
tanks have been completed successfully,

AGINA "5"
XLRS1-Ba-7 Engine

© Preliminary Flight Rating tests were completed
on 18 February and a formal review of the test
results has been started. Both of the engines being
tested experienced thrust chamber burnout problems
during continuous 240 second firings under certain
conditions of coolant ambien} temperatures and
IRFNA solids content. Development efforts are being
continued by the engine manufacturer to alleviate
these problems. In the meantime, however, satisfac-
tory chamber cooling can be assured by closely con-
trolling the solids content and the ambient temper-
Oture (70 degrees maximum) of the oxidizer used,
4

©@ Testing of the XLR81-Ba-7 engine installed in the
Propulsion Vehicle Test Assembly was completed at
Santa Cruz Test Base. The total testing program
consisted of seven firings, three of which were con-
ducted during February. Two Engine restart oper-
ations were performed successfully in this test series.
Testing was accomplished on a propulsion test
vehicle assembly having aluminum tanks and ‘a test

installation of the fuet pressure driven hydraulic con-
trol system (See DISCOVERER Monthly Progress sec-
tion}. On 8 February a full duration run of 244
seconds was made. On 9 February, the engine was
tun for 220 seconds, shut down for five minutes, and
restarted for a 28 second firing to oxidizer exhaus-
tion, On 12 February, the engine was operated for
220 seconds, shut down for five minutes, and re-
ignited for a 27 second firing.

AGENA “B""
XLRB1-Ba-9 Engine

© A test run of the XLR81-Ba-9 ‘engine installed in
the propulsion test vehicle assembly was made on 19
Febraury at the Santa Cruz Test Base. The engine
was operated initially for 120 seconds, shut down
for five minutes, and reignited for a 122 second
Operation. The first successful gimballing program
was achieved on this run. Six additional test firings
.Br6 scheduled for this engine,

© “Testing of the ‘engine nozzle extension Program
[to achieve a 45:1 area ratio} was continued during
the month at Bell Alrcraft Co. An extension con-
structed of titanium was tested successfully for 240
seconds, followed by a 10 second restart firing,
Another successful 240 second firing was" acconi-
plished on the same extension following a recoating
of the interior surface with cluminum oxide.

13


Figure 1. Artist's concept of MIDAS
satellite (righ). Diewtnts baow how

THERMA! TM EXIP—EQUIP~OXIDIZER~FUEL

SHIELD“ ANTENNA SEPARATION ANTENNA COMPT. TANK TANK
oodles th GENA : vegan = 18,100

coatiguretion except je zi
‘or soor vont Fool, RP-1 . 74,900
‘Oxtdizer (LOX) 172,300
GROSS WEIGHT (fbs.) 262,300
. Engine—MA-2

‘Thrust {Ibs, vac) Boost 356,000

Susteiner 82,100
‘Spec. Imp. (sec. voc.) Boost 26

won ie |

Sustciner 310



“CONFIDENTIAL

fully. Test results indicate that the propellant bonding
“fix” has solved the burn-through problem and Pro-
vide high confidence in the success of two simulated

@ ANASA Sergeant type launcher was procured
from Vought Asironoutics for lounch of Project 609A.
unguided vehicles,

Payload Priorities Committee Formed

©@ A committee has been formed to review, vall-
date, and assign Priorities to payload experiments
Propesed by ARDC Centers for Project 609A and
other test support Programs, such as ATLAS Piggy-
back. The committee includes, representatives from
various AFBMD directorates of WDZ and WOTV. A
charter and set of Operating rules are being

Prepared,

WDLPM-4 188


“CONFIDENTIAL

by long range plonning Gnd maximum integration
with other programs. Use of the basic 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 sta-
bilizing the fourth stage without spin and use of the
vehicle in less than the full ‘four-stage configuration,
Close integration with the current ballistic missile
Program will effect an economy by permitting tests
and experiments to be conducted on regularly sched-
vled ballistic missile test flights whenever possible
without delaying schedules, Economy in the opera-
Honal 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. RELIABILITY will
be obtained by a nine of ten vehicle R&D flight test
Program, at least four flights of the basic SCOuT,
and maximum use of ki gained in prior Air
Force ballistic missile flight testing. VERSATILITY will
be achieved by designing a vehicle capable of being
readily adapted fo a wide range of payload varia-
flons, and capable of being flown in several configu
rations of four stages or less, This VERSATILITY resutts
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 @pproxi-

ua "CONFIDENTIAL

mate maximum of 400 miles with 150 Pound pay-

Program Management—An abbreviated devel-
opment plan, covering the R&D phase only, was
@pproved on 9 January, 1959. Funds in the amount
of $8,180,000 were made available for this abbre-
viated portion of the Program only. A letter was issued
assigning management responsibility to AFMBD, with
emphasis on integrating the program with the scien:
tific and military research experiments conducted on
regularly scheduled ballistic missile flight tests (Piggy-
back Program). In June 1959, Aeronutronic Division
Of the Ford Motor Company was chosen through nor-
mal competitive bidding as the Payload, Test, and
Systems integration Contractor, Arrangements have
been made for the procurement of vehicle compo-
nents and associated support equipment, modified
to meet Program 609A requirements, through NASA,
rather thon through the SCOUT Program contractors.
Atlantic Missile Range facilities consisting of launch
complex 18 will be made available fo the Air Force
for this program. A Project 609A division has been
established within the 6555th Test Wing (Develop-

ticipate in the ‘assembly, checkout and launch opera-
tions of the R&D phase under the direction of the °
Payload and Test Contractor. An all-military opera-
tional capability will be developed from within this
group.

Figure 2.
Four diferent .
secerns
possible using
different

arrangements
of Project GO9A
sages.

WDLPM4 188


WDLPM4 1g9 ©

CONFIDENTIAL

FIBERGLAS
ATTACH RING

HEAT
SHIELD

ALY

PERISCOPE & UMBILICAL DOOR

~ FLOTATION
BAGS ACCESS ‘AF TERBODY

Figure 3. McDonnell shell so be carried on second
ATLAS/MERGURY flight test,


MONTHLY PROGRESS— COMMUNICATIONS .
SATELLITE Program
Program Administration

e AFBMD has been advised by OSD teletype that
Amendment 4 to ARPA Order No: 54, dated 29 Feb-
‘vary 1960, has directed that efforts defined under
Projects STEER and DECREE be reoriented into a
single integrated Project under the code name
ADVENT.

@ AFBMD is Preparing a development pian which
reorients the program in keeping with the spirit and
intent of Amendment 4. The new plan will be sub-
mitted to ARPA in March.

e Because FY 1960 funds have not been released,
current launch schedules for Projects STEER and
DECREE are not valid. As a result, the efforts of
Project STEER contractors Were reoriented during
February to preserve the integrity of the engineering
team as long as possible.

Technical Progress

© Preliminary configuration studies of the final
stage vehicle have been completed. These include
detailed weight, Power, thermal analysis and volume
estimates. Paris counts were made on system block
diagrams to provide realistic weight and power
estimates, Super insulation is used on surfaces
exposed fo solar radiation, and variable tempera-
ture contro! radiators on surfaces not $0 exposed,
These studies were conducted by Space Technology
Laboratories.

@ Laboratory tests of components and mock-up
models have been performed by the General Electric
Company, emphasizing the following areas:

@. Tests of siiver-zinc batteries for Possible use as
© primary power -source during periods of eclipse.
Test failures hove justified a decision to use nickel-
cadmium batteries,

b, Radiation testing of solar cells,

. Communications antenna Pattern tests on full-
scale model.

d. Telemetry antenna Patern tests on quarter-
scale model.

@. Bearing tests under vacuum with various com-
binations of teflon and silver-plated steel! balls,

f. Emissivity evaluation tests on several materials
9s part of the temperature contro! method.

@ Extensive analyses relating to the development
of a communications satellite have been formalized

by Space Technology Laboratories and transmitted
to General Electric, Areas covered include:

@. Station keeping accuracies,

b. Visibility time.

¢. Use of elliptical orbits for Polar coverage.

d. Trajectory error budgets.

. Satellite space radiation environment.

f. Vacuum bearing dota,

g. Electrical power and weight trade-offs.

h. Various satellite attitude contro! systems and
contro! laws,

i, Separable. versus hon-separable final stage

icles.

|. All available ABLE Program antenna design
data.

@ Studies currently in progress which affect the
design of the final stage vehicle include:

a. Analysis of problems associated with orbit sta-
tion keeping to determine the required accuracies
for command and mechanization,

b. Analog and digital simulations to evaluate the
effects of lunar and solar gravitational Perturbations,
gravity gradient and solar radiation pressure.

¢. Basic battery studies, ie, reliability versus
cyclic life and depth of discharge on voltage tegu-
lation.

d. Effects of radiation and micrometeorite impact
on solar cell efficiency.

©. Several proputsion methods for the final stage
vehicle, including the use of bi-propellants and
mono-propetlants.

Communications Equipment Development
UHF .

© Bendix has conducted an experimental evalua-
tion of hard limiters for use in the satellite repeater
‘equipment. Their evaluation exactly confirms theo-
retical predictions and has shown the following
effects of hard limiting upon frequency modulated
or phase modulated signals, For positive or negative
signal-to-noise or signol-to-jamming ratios of greater
magnitude than about 10 decibels, hard limiting
increases the magnitude of the ratio by six decibels.
Thus, in the absence of jamming the hard timiter
would act as a noise filter; however, in the presence
of overpowering jamming (negatives $/J ratios) the
hard limiting would enhance the fomming signal and
reduce the intelligence-carrying signal.

"= {tie WDLPM4 188



H2

1

2.

10.
ve
12.
13.
14,

16.

17,

18.

TABLE |. TIROS POWERED TRAJECTORY FLIGHT PLAN

” Vertical lift-off

Stage | pitch program begins

BIL guidance begins Stage | closed-loop steering
End Stage | pitch program; end Stage | steering

Stage | constant attitude flight

MECO—Stage | main-engine cutoff

3.8 sec of vernier operation prior to Stage I-II separation

Stage I} separation and ignition; begin Stage II pitch program

BTL guidance begins Stage I! closed-loop steering

Jettison nose fairing

BTL guidance ends Stage II closed-loop steering
BTL discrete ends Stage Il pitch program

BTL discrete spins up third stage and payload—120 rpm

Stage II cutaff—SECO; begin coast period

Stage {I-lll separation

Fire Stage Ill rocket

Separate payload from Stage Ill
De-spin payload—12 rpm

pens .
ry

10 sec.

90 sec.

130 sec.
130-159 sec.
159 sec approximately.
159-162.8 sec.
162.8 sec.
172.8 sec.
182.8 sec.

253 sec.

255 sec.

265 sec.

SECO (Approximately 267 sec.)

268.5 sec.
SECO—2 sec.

811 sec to
848 sec.

864 sec.
in orbit

WDLPM-4 188


! AQINA —AGENA
prt ope
| 13701400
6590 13,100
a 100
. 7987 14,800
' YUNIeS XLBI-t07
7-— FUEL AND
brad ‘5g00 OXIDIZER
120 40 TANK
Ne Yes SECTION
SM-68 DMT
6.950 5.950
7 3370 33750
, £8900 _ 68,300 ‘SEPARATION
109,000 “108,000 PLANE
mes a3
Block 1 Block 2

SPHERE. ENGINE (UDMH)
Figure 1.” Photograph oj swo-stage DISCOVERER vebicle (left) and

| ~ detailed drawing of AGENA, second stage (right),

AGENA POWER

GAS VETS (ROLL)
GIMBALLED ENGINE
VELOCITY METER
————— <>

~

MMAND
Time TO FIRE
t——— COMMAND 6
VELOCITY TO

BE GAINED
|~> transmitter



Program Objectives

1. Provide accurate navigational reference informa-
fion for POLARIS launches.

2. Precise determination of satellite position by
measuring the doppler shift of satellite transmitted
radio signals.

3. Investigate the refractive effect of the jonosphere
on radio transmissions.

4. Acquire additional geodetic and geographical
data by precision tracking of the orbiting satellite.

Flight Vehicles ‘TRANSIT 1A consisted of three
stages as shown in Figure » TRANSIT 1B, 2A
and 2B are two-stge vehicles as shown in Figure.

Launch Plans Alt vehicles will be launched from
Atlantic Missile Range pad 17A or 17B. Launch
azimuth for TRANSITS 1A and 1B Is 44.5 degrees
and for TRANSITS 2A and 2B, 140 degrees.

marsree A
F2 OEE WDLPMA 188

8 SECOND STAGE—AJ10-42 (Aerejet-General)
Three at altitude 7700 pounds
Ty Specific impulse (vac) 271 seconds
Total impulse (min) 870,000 Ibs/s0c
Burning time 115 seconds
Propellant Uquid
H FIRST STAGE—THOR IRBM
Thrst (a) 151,500 pounds
Specific Impulse (1.1) 248 seconds
Specific impulse (vox) 287 seconds
Hy Burning time 158 seconds"
Propeltont Uquid
A '
t
1 —TRANSIT 1B, 2A and 2B

UPPER PAYLOAD*
PAYLOAD

(36-INCH DIAMETER)
PAYLOAD

SEPERATION PLANE

PAYLOAD
SPIN MECHANISM

EQUIPMENT
COMPARTMENT

PROPULSION SYSTEM
54.79 DIAMETER
GIMBAL ¢

ISgtace Iv
SEPARATION PL ANE

STAGE I/IL ADAPTER
STAGE I

%* 20 INCH DIAMETER
TRANSIT 2A & 2B ONLY

Powered Right Trajectory The powered flight
trajectory for TRANSITS 1B, 2A and 2B ‘is shown and
described in Figure « The sequence of events
from launch through payload separation for TRAN-
SIT 1B is given in Table 1.

Payload Description The spherical payloads are
approximately 36 inches in diameter and weigh
between 200 and 270 pounds. Payload equipment
includes four transmitters (on frequencies of 54, 108,
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. The TRANSIT 1B
payload will also contain an infrared scanner which
will operate for the.first four days of orbit. On TRAN-
SITS 2A and 28 a 20-inch sphere, mounted on top
of the 36-inch sphere, will contain instrumentation
for studying solar emissions. The payloads will be
spin-stabilized in orbit. .

GUIDANCE COMPARTMENT



The Communications Satellite Program will investi-
gate, in two phases of increasing complexity, the
feasibility of using synchronously spaced satellites as
instantaneous fepeaters for radio communications.
Under ARPA Order No. 54 as amended, AFBMD is
responsible for design, development and test of the
complete system, including launch, satellite tracking
and control, and fecessary support facilities and
units, Wright Air Development Division is responsible
for the development of aircraft communications
equipment for both phases. Responsibility for sat-
ellite and ground station communications equipment
is assigned to WADD {first phase) and to the Army

SUN TRACKERS

showing antenna and paddle orienta-

bi, tp Communications segllive | (STi ER Phase)
Small drawing shows operational conceps of
ie.

IR SCANNER

PHOTOVOLTAIC CELLS ~~

Signal Research and Development Laboratory
(second phase). The two phases of the program have
been designated STEER and DECREE. The description
and objectives of each phase are as follows:

‘STEER (Figure 1)

This four-Aight test phase will use ATLAS/AGENA.
vehicles to inject satellites into polar orbits with six
hour orbital periods. This phase stresses earliest pos-
sible availability consistent with program objectives.

~ COMMUNICATION ANTENNA

TRACKING/ TELEMETRY
ANTENNAS (4)

NOZZLES

Communications
Satellite

Dl



Figure 1. ABLE-3 fight sest vebicle
(left) of four-stage ABLE-3 vebicle,

brine launched from Atlantic Missle Range. Dimentional deaving

El


WDLPM-4 188

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 preampli-
fers will be used because of the low-level signal to
be expected from the aircraft transmitters.

Aircraft communications equipment will make maxi-
mum use of UHF equipment presently installed on SAC
aircraft. Miner modifications will provide an ade-
quate signal-to-noise ratio from the satellite trans-
mitter at a maximum range of 8,000 nautical miles,
Three different approaches fo 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
teceiver transmitter is being developed which would
teplace 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
transmitters. This unit will be capable of the second-
ary function of providing @ certain amount of jam-
ming power in the UHF band when not in use as a
communications transmitter. Should spread-spectrum
modulation be required to overcome interference or

(SEC) (FEET)

TIME ALTITUDE VELOCITY
(FEET/SEC)

— —_ —_
BOOSTER SEPARATION 139 208,000

9,400

SUSTAINER SEPARATION 295 566,000

21,000

STAGE II IGNITION 297 567,000
pia

21,000

STAGE Il Ist CUTOFF 395 570,000

30,800

STAGE II 2ND IGNITION 6694 34x10
| (ORBIT INJECTION PHASE)

12,100

STAGE II FINAL CUTOFF 6667 34106

16,000

SATELLITE SEPARATION 6677 34x 108

16,000

jamming on the aircratt-to-satellite link, compatible
modifications would be made to aircraft equipment.

GROUND SUPPORT FACILITIES — Ground tracking
and data handling capability is required to: (a) Ver-
ify that the satellite has been injected into orbit,
{b) provide data on performance of the final stage
vehicle in orbit, and (c) provide sufficient control to
Permit synchronization of the satellite Position in
relation to other satellites in the total system. Inves-
tigation is being made of the possibility of using
SAMOS and MIDAS Program ground support facil-
ities. Ground stations at Offut 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.

DECREE

The ten flight tests in this phase will provide the
R&D effort for demonsirating the feasibility of a
global communications system, using precisely
spaced “hovering” satellites which, essentially, have
orbital periods of 24 hours.

heloal.! bat
“oF BENTAL


Flight Time Inertial Speed Downrange Altitude
Point {seconds} Comments (tt/seck Distance {n.mJ) (namJd
2A fe ats 18 2A 1B 2A 1A 2A
“18 2.5 _ ‘Tronsit vehicles 23 2B 23
1 w 10 End of vertical rise 1,346 1,346 0 ° 0.077 0.077
2 167 167 __ First stage burnout 13,611 12,929 75.2 797 AQ 483
End of second stage
3 442 Best bu ring period 24,539 24,376 785.6 7780 200.1 203.0
4 1489 1a Resto second tloge aa 46 72599 428240000 ©5000 «5000
5 1,504 1,462 _ Injection into orbit 24,258 24,259 44163 4130.0 500.0 500.0

FLIGHT TRAJECTORY —TRANSIT 18, 2A ond 28

i

Orbital Performance Achievement of program
objectives is based primarily on measuring the
doppler shift of satellite transmitted radio signals.
During the first three months of flight, the four trans-
mitters will be operated to obtain experimental con-
firmation of the theoretical mathematical relationship
between the frequency and the refractive index of
the ionosphere. Studies have shown that refraction
effects on the doppler shift can be eliminated by
using the transmission from two satellites. After four
months of tracking the satellite by measuring the
doppler shift of the satellite radio signal, the exact
Position of the satelilte at any point in the orbit
should be known. Using known orbital positions,

_WDLPM-4 188°

ween vr

ships and aircraft can then use satellite signals to
make anclogous computations to establish accurate
position. Navigational fixes of 0.1 mile accuracy are
expected to be obtained.

Ground Support Stations Tracking stations will
be operated in Maryland, Texas, New Mexico, Wash-
ington and Newfoundland. First and second stage
tracking and telemetry and second stage guidance
will be provided by the Atlantic Missile Range. A
mobile tracking and telemetry van will be located in
Germany for TRANSIT 1B and South America for
TRANSITS 2A and 28. These locations were selected
@s the closest sites possible to the orbit injection
point.

F3


WDLPM-4 188

MONTHLY PROGRESS —MIDAS Program

"Hight Test Progress

©@ The first MIDAS flight test vehicle was launched
from Atlantic Missile Range launch pad 14 on 25
February. Satellite orbi ined.
talled analysis of the flight and of the problems en-
countered will be included in next month's report.

Flight Parameters

@ The first MIDAS flight test was Programmed to
place the satellite vehicle into a 261 nautical mile
circular orbit, with a maximum eccentricity of 0.007
and an inclination angle (to the equator) of 32.5
degrees. A useful orbital lifetime of 29 days was
anticipated. A launch azimuth of 107 degrees was
used, with orbital injection planned to occur at T
plus 655 seconds at a velocity of 25,024 feet per
second.

Launch Preparations

e The electrical rewiring of launch pod 14 re-
quired for the MIDAS vehicle, and launch pad um-

bilical drop tests, were completed on schedule. Addi-
tional redundant electrical circuits were installed in
the umbilical mast to provide. increased launch re-
liability.

© Systems checkout of the ground support equip-
ment was conducted successfully with no problems
becoming apparaent. Checkout of the ATLAS
booster also was conducted with completely satis-
factory results.

@ MIDAS vehicle simulators were delivered to
AMR, Kaena Point (Haweil) and Vandenberg AFB
early in February. These units were used to train and
familiarize operating personnel in vehicle handling,
checkout, tracking and readout; and for electrical
checkout of associated ground equipment. Each unit
consists of two equipment racks. The simulators are
capable of receiving telemetry, transmitting com-
mands, and simutating the characteristics of the
infrared payload and communications subsystem of
the orbiting MIDAS satellite,

Satellite Readout Pians

@ AMR, Kaena Point and Vandenberg AFB were
scheduled fo perform payload-to-ground dota link
readout. All three stations were to have tape re-
corded the satellite system-time data for analysis
and processing for presentation on the command
console of the ground presentation unit at the Satel-
lite Test Center, Sunnyvale, California. in addition,
real-time readout was to have been performed on
the ground presentation unit at Vandenberg AFB.
Motion pictures were to have been made of the real-
time ground presentation, with comparable system-
time indicated on each frame.

e A series of targets had been planned to test the
infrared readout capability of the orbiting MIDAS
satellite. These inclsded the launchings of an ATLAS
and a TITAN missile from the Atlantic Missile Range,
and the SAC founch of an ATLAS missile from Van-
denberg AFB. All launches were to have been timed
to coordinate with passes of the MIDAS satellite. in

" addition, ten pyrotechnic targets were to have been

ignited at Vandenberg AFB and Edwards AFB during
night time orbital passes.

Technical Progress

Second Stage Vehicles

® Preparation of AGENA vehicle 1007 for instal-
lation on the second MIDAS flight vehicle is pro-
ceeding on schedule in AMR Hangar E. X-ray exam-

Figure 6. View of ATLAS booster for frst MIDAS flight,
a8 seen from gantry. ATLAS is installed in transporter prior
40 erection.


Figare 6. AGENA "” vebicle

6. ile shown mounted
Center. This is the fual AGENA "A Soe

Bilomedicat Recovery Capsule

© Tests of the biomedical capsule designed for a
small primate were resumed in the Lockheed high
altitude temperature simulation chamber on 8 Feb-
tuary. The General Electric capsule tested utilized
several modifications and techniques derived from
thermal profile tests in November and proof tests by
the School of Aviation Medicine in December. These
include increased cooling capacity, refinement of
sensor-to-animal attachment methods for telemetry
readout, relocation of life chamber components, and
reprogramming of psychomotor response stimuli. The
first full-duration test of the capsule containing a live
peimate was completed on February 12. This 55-hour

teat_simulating 4 complete flight wos initiated at

WDLPM-4 188

im trensporter-erector at Modification and Checkout
‘thadaled Jor las program (OSC Oy ee etsEaeM

__Yandenberg AFB with the primate sealed in the cap-

hours the capsule was flown to the Sunnyvale Devel-
opment Center (with Passenger). It was placed in the
High Altitude Temperature Simulator for the simu-
lated orbital phese, then sequenced through simu-
lated re-entry-recovery phases. It was then placed in
temperature-regulated water for five hours to simu-
late the final five-hour recovery phase. During these
tests the primate responded to stimuli properly and
was able to perform all programmed tasks. A new
feeder, designed by the Schoo! of Aviation Medi-
cine, proved excellent. Electrocardiogram readouts
were excellent and all components of the air regen-
eration system functioned well.

Performed, “and after’ 22°

A?


‘WDLPM-4 188

-CONFIDENTIAL

Computer and Guidance

@ All guidance constants were computed for the
launch period and verified by making a series of
flight simulations on an IBM 709 computer, using a
dosed loop guidance computer program. A new
wiring list and answer tapes and plots were gener-
ated on the IBM 709 for use with the Burroughs
(J-1) Guidance Computer ot AMR. Flight simulations
on the J-T computer produced output tapes and plots
which were in complete agreement with those gen-
erated on the IBM 709.

Antennas

© Life tests were conducted on the Transco Coaxial
Relay at pressures simulating altitudes of 385,000
to 415,000 feet. With 15 volts applied to the sole-
noid, the switch was cycled 4,500 times with no
malfunctions. Centrifuge and vibration tests also
were completed successfully with 10 voits applied to
the solenoid.

Ground Support Stations

@ At the end of February, all ground stations were
ready to support the ABLE-4 THOR launch. These
include: AMR, Manchester. {England}, Hawaii, and
Singapore.

ABLE-3
Phase Comparison Data

© Reduction analysis of satellite phase comparison
data commenced in February. The reduction had
been awaiting computer runs of the expected values
of the angle formed by the planes containing (a) the
vehicle spin axis and sun vector and (b) the spin
axis of magnetic field vector. This data will be used
fo test for primary winds in the outer atmosphere.
Primary winds would be expected to produce a cur-
tent which would result in a magnetic field perturba-
tion.

Magnetic Radiation

@ Two scientific papers {originated in Space Tech-
nology Laboratories} were presented at the New

_ York meeting of the American Physical Society, The

most significant and scientifically rewarding result
described in these papers is the observation of the
behavior of the second Van Allen radiation zone
when large flares occurred on the sun. A sudden
magnetic storm, presumably of solar origin, was
observed on the earth between 15-16 August 1959.
Fluctuations in intensity, sometimes rising to 5,000
times the normal base value in a matter of a few
seconds, were observed in the outer Van Allen zone
on 16-17 August. Simultaneously, the boundary of
the outer zone moved in towards the earth by almost
4,500 miles. The approximate boundaries of the
radiation fields were determined from data the satel-
lites had been sending for over o week. Between
20-29 August, the boundaries moved out to a dis-
fonce even further from earth than their ante-storm
position.

@ These results show that during magnetic storms
particles are thrown off from the second radiation
zone, presumably by a fluctuating magnetic field.
Earlier evidence of a fluctuating magnetic field had
been obtained from a magnetometer carried aboard
Pioneer | and Explorer IV, and also from the intensity
fluctuations observed by Explorer VI's scintillation
counter.

@ After the storm, particles began to fill the second
radiation zone; intensity increased and the bound-
aries moved further from the earth.

Electron Flux

@ Other results from Explorer Vi include observa- °

tion of flux of particles thought to be electrons on
the order of 100,000,000: electrons/cm* sec, and
the first detailed mapping of the isointensity contours
in the outer radiation zone. The latter revealed a
Pronounced concentration of particles around the
geometric equator at high altitudes.

E7



formed as planned over its entire powered flight
trajectory. The frajectory of this flight, from the
Atlantic Missile Range to the vicinity of the moon,
was established to achieve the tightest possible cir-

ABLE-£ THOR is a deep space probe which was
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 fiquid fueled,
guided second stage, and ABL248A.3 solid-fuel
third stage, This vehicle will place a 90 Pound pay-
load into space flight. The payload ‘contains no retro-
rocket, and no. Gitempt will be made to intercept
Venus. The design of the Payload components is the
same Gs those which have been proven on. the
ABLE-3 satellite.

WDLPM-4 188


(d) Testing of techniques for recovery of a capsule
ejected from the orbiting satellite,
{e) Testing of ground support equipment and develop-
+ Ment of personne! Proficiency,
"(0 Conducting bio-medical experiments with mice and

small primates, including injection into orbit, re-
entry and recovery.

Early tests confirmed vehicle flight and satellite orbit
capabilities, developed system reliabitty and predict-
ability, and established ground support, tracking, and
data acquisition requirements, Subsequent flights are
planned to acquire scientific data for design of
advanced military reconnaissance Payload compon-
ents. Typical data gathering objectives include: cosmic
and atomic radiation, magnetic field, total electron
density, auroral radiation, micrometeorite measure-
‘ment, Lyman alpha from space (or stars), solar radia-
tion, and atmosphere density (drag) and composition.

A4

separation point, and impact ares, Inset at left shout
vebicle attitude required to effect capsule re-entry
‘ollowhee vious

A world-wide network of control, tracking, and data
acquisition stations has been established. Overall oper-
ational control is exercised by the Control Center in
Palo Alto, California, Blockhouse and launch ‘opera-
tions are performed at the Vandenberg Air Force Base
Control Center.

Telemetry ships cre positioned @s required by the
specific mission of each flight. Figures 2 and 3 show
@ typical launch trajectory from Vandenberg Air
Force Base, and figure 3 shows schematically « typ-
ical orbit. An additional objective of this program is
the development of @ controlled re-entry and
recovery capability for the Payload capsule {Figure
4). An impact area has been established near the
Hawaiion Islands, and a Tecovery force activated.
Techniques have.been developed for cerial recovery
by C-119 circraft and for sea recovery by Navy
surface vessels. The recovery phase of the program
has provided advances in re-entry vehicle technol-
ogy. This information will be used in support of more
advanced projects, including the return of a manned
satellite from orbit.

DATA CABLE

16 th ORBIT



Figure 3. Typical DISCOVERER trajectory (sbooe)

from launching at Vandenberg APB 40 orbit. Typical

satellite orbital path around the earth (right),

The DISCOVERER Program consists of the design,
development and flight testing of 29 two-stage vehi-
cles (Figure 1), using the THOR IRBM as o first stage
booster and the AGENA vehicle, powered by the
Bell LRE? rocket engine series as the second stage
satellite. The DISCOVERER Program was established
early in 1958 under direction of the Advanced
Research Projects Agency, with technical manage-
ment assigned to APBMD. On 14 November 1959,
Program responsibility was transferred from ARPA
to the Air Force by the Secretary of Defense. Prime
contractor for the program is Lockheed Missile and
Space Division. The DISCOVERER Program will Ppro-
vide: (a) space research in support of the advanced
military reconnaissance satellite systems, Programs,
{b) test of the ground communications and tracking
network for these programs, and (e) flight testing
of the AGENA second stage vehicle.

PALO ALTON
VAMDENSEROAQ

Primary objectives include:

(a) Fight test of the satellite vehicle airframe, pro-
pulsion, guidance and control systems, auxiliary
Power supply, and telemetry, tracking and com-
mand equipment,

(b) Attaining satellite stabilization in orbit,

(c) Obtaining satellite internal thermal environment
dota,



View of launch pad 14 gasiry and umbilical rower
erection of ATLAS booster for first MIDAS Highs.

8. AGENA sebicle (above) mounted in transporter
base (right) being lifted into gantry
ber.

ination of the thrust chamber revecied the Presence
of a foreign particle in the oxidizer inlet manifold.
A new thrust chamber was Installed on 8 February.

@ Design of the AGENA vehicle for the third
MIDAS flight test is Proceeding on schedule. Struc-
tures are being fabricated and release of equipment
and installation bracketry designs to manufacturing
is anticipated early in April.

Infrared Scanner Units

@ Three of the infrared scanner units for the first
two MIDAS flights were shipped to AMR during
February, and the fourth is in the Modification and
Checkout Center. One of the units at AMR is the
flight article for the second MIDAS flight.



AS

MONTHLY PROGRESS—.DISCOVERER
. Program ‘

Flight Test Progress

DISCOVERER IX

@ DISCOVERER IX was launched from Vandenberg
AFB pad 4 at 1051 PST on 4 February. No problems
were encountered during the final countdown. Liftoff
was normal except for a malfunction of the helium

quick disconnect. Alth

ough the initial ascent portion

of the flight appeared formal, instrumentation indi-
cated early termination of both THOR and AGENA
burn periods. Data analysis revealed that the THOR
engine shut down Approximately 19 seconds early,
resulting ina velocity loss of about 4,000 feet per
second. Premature shut down of the AGENA engine

tesulted from the loss

of propellant tank pressuriza-

tion caused by the helium quick disconnect matfunc-

tion at liftoff. AGENA

impact occurred in the ocean

about 400 miles south of the launch site.

Quick Disconnect Problem

+ @ At the time of DISCOVERER IX liftoff, the helium
quick disconnect coupling released 1.1 second late.
During this delay, the release lonyord was broken,
resulting in a pull on the coupling. The vehicle
mounting bracket broke, allowing the airborne half

of the coupling to be

Pulled out of the vehicle. The

coupling was recovered and has been subjected

inside the coupling whi

Drain holes have been

ich blocked the release action.
Incorporated to Prevent a

recurrence of the problem.

DISCOVERER X

e DISCOVERER X was launched from Vandetberg
AFB pad 5 at 1215 PST on 19 February. The count-

down proceeded smo

‘othty and launch was accom-

plished on the first attempt. immediately after liftoff,
THOR booster pitch oscillations began and, at T plus

56.4 seconds the vehi

ice was destroyed at 20,900

feet by the Range Safety Officer. Both the THOR and
AGENA destruct systems operated satisfactorily.
Many major Components were recovered and are
being examined and analyzed. No personnel injury
Or property damage was sustained. Preliminary
analysis indicated that a THOR autopilot malfunction
caused the main and vernier engines to oscillate
between hardover stop positions, starting at liftoff.
Extensive studies are underway to ascertain and cor-
tect the responsible conditions.

Flight Schedule

© DISCOVERER XI is scheduled for faunch from

Vandenberg AFB in
carry an advanced en;
instrumented recovery

late March. This vehicle will
gineering test payload and on
capsule.

Technical Progress
Second Stage Vehicles

@ All of the remaining AGENA “A” vehicles (DIS-
COVERER flights Xi through XVI} are ot Vandenberg
AFB in various stages of launch preparation. Three

AGENA Propulsion System
See AGENA Monthly Progress section.

Guidance and Control

© A hydraulic control system driven by fuel pres-
sure is being developed to save weight and electrical
power. The system replaces the electric motor driven
hydraulic pump used on AGENA "A" vehicles and is
being tested currently at Santa Cruz Test | Boer A
weight saving of approximately 20 poun: will be
tealized. The system is plonned for incorporation on
DISCOVERER 21. This flight will be the first to use
the improved configuration of the AGENA “B"
{XLR81-Ba-9 engine and thrust chamber extension).

Figure 3. Optical wracking lights installed ipment
rath of DISCOVERER TS AE BN eid ct Ht eesibmens



WOLPM-4. 188

a ils

© Site selection criteria for the communications
subsystems ground station have been prepared. No
Sction hos been taken because of a lack of funding.

Microwave

@ Preparation of Specifications for satellite borne
end ground station communications equipment is
nearing completion. Consideration has been given
to the terminal ‘equipment for two ground stations.
The terminal equipment will consist of: 8 AN/FGC-
25X teletypewriter sets, 4 AN/TCC-41 PCM multiplex
units, 4 AN/GXC-4 facsimile units, 8 AN/TCC-30
terminal units and 4 ‘analog-to-digital converters.

©. Army Signal Research ond Development Labor-
atory has requested the following specific frequency
allocations for the communications subsystem:

Ground station-to-satellite—7140 me, 7580 mc,
8020 me, 8460 me.

Satellite-to-ground station — 1830 me, 1940 me,
2050 me, 2160 me.

Beacon tracking of satellite— 1770 mc, 1880 mc,
1990 me, 2100 me.
@ A contract was awarded Varian Associates for
the: purchase of one high power, 8000 mc klystron

tube and the performance data to indicate the sys-
tem suitability of this tube,


a foreword fOcee

a

eres

- WDAT 60-1495


MANNED SATELLITE

RETRO & POSIGRADE ROCKETS.

ROLL JETS EAT SHIELD

ENTRANCE HATCH . \

scare’

ESCAPE ROCKET-
JETTISON il

ESCAPE ROCKET

HORIZON SCANNER
PITCH & YAW JETS

ROCKET OSIGRADE.
RE TROGRADE-
WEIGHT AT SEPARATION APPROX 25 LBS ORBIT INCLINATION 33 DEGREES
ORBITAL ALTITUDE 105-115 MILES (n) HEAT SHIELD ABLATIVE
ORBITAL CYCLES 3-18 RECOVERY WATER OR LAND

Figure 1. Complete vebicle (top view) with satellite installed ow ATLAS booster. Manned satellite
(bottoms view) showing pilot’ ight position, and detail views of retro and potigrade rockets eed piles
safety system escape rockets,

WDLPMA 188 CONFIDENTIAL 4


MONTHLY PROGRESS—SAMOS Program
Program Administration

@ Asa result of program reorientation announced
during December 1959 and made necessary by re-
stricted funding levels, the flight test plan for ferret
Payloads has been revised as follows: .

Number of Flights
Payload Old Schedule New Schedule
Fl 2 3
F.2 2{0—F-2A) 4
(1-28)
3 A(2—F-3A) 1
(2—F-38)

© A combined visual/ferret payload will be tested
on the first 3 flights. The first seven ferret payloads
(F-1 and F.2} will include Progressively more com-
plete installations of receivers and antennas to pro-
vide increasingly greater electronic measurement
capability. The major portion of the hardware com-
ponents developed for the original program are
usable in the reoriented program.

Technical Progress
Second Stage Vehicles

@ Work on the second stage (AGENA) vehicle for
the first SAMOS flight is 70 percent complete in the
Modification and Checkout Center. This vehicle will
be the first of three to carry a combination visual
and ferret (E-1 and F-1) payload. Assembly of the
other two vehicles is proceeding on schedule. In-
terior design of the AGENA vehicles for flights 4 and
subsequent is proceeding on schedule. A common
airframe design from the forward equipment com-
partment aft is being used for these vehicles and for
MIDAS vehicles (flights 3 and subsequent). Equipment
installations need not be interchangeable. Substan-
tial progress has been made on the design of the
AGENA vehicle to be used for E+5 (recoverable)
SAMOS payloads.

Visual Reconnaissance System

Viswal Reconoaisscace System payloads ore being devel-
oped ino minimum number of configuration: to attain
readout and recovery objectives. The detigaction ond
purpore of each configuration Is os follows:

Readout:
E-1 — Component Test Payloads
£-2 — Steerable Reconnaissance Pay-

loads (with 20-foot ground reso-
lution)

. Recovery:
E-5—High Resolution, Recoverable
Payload {with 5-foot ground
resolution)

Payloads

@ E-1 Payloads—The first E-1 flight article pay-
load was delivered to LMSD on 8 February. Func-
tional tests were performed on all components. Dur-
ing a preliminary functional test, with the payload
mounted in the collimator, a system. resolution of
greater than 94 lines per millimeter was obtained.
The payload was subjected to o series of three 19-
hour tests under simulated. orbital conditions, with
satisfactory results being obtained. The second &1
Payload is undergoing. quality evaluation testing at
Eastman Kodak. This is a spare Payload for compon-
ent replacement only and-will be delivered to LMSD
unassembled before 15 March.

@ = E-2 Payloads—Delivery of the first E-2 payload
is scheduled for July. Environmental tests of the ther-
mal model E-2 payload were completed on 28
January in the high altitude temperature simulation
chamber. Test objectives were achieved. Changes in
the payload housing surface and heater power re-
quirements are being made as a result of testing
data obtained.

@ £-5 Payloads—Design of the high acuity pano-
ramic camera system is Proceeding satisfactorily.
The special opticat gloss for the lens elements, which
has been ordered from West Germany, will be de-
liverd to the tek Corporation in mid-April, The De-
velopment Test Pian for the recovery capsule hos
been published, including Payload test requirements
from checkout through post-launch operations. Avco
Corporation is conducting wind tunnel tests on vari-
ous capsule configurations as a parallel effort with
LASD aerodynamics studies.

Ground Support Equipment

©@ The complete visual reconnaissance system
ground support equipment complex was operated
with the E-1 payload during February. All equipment
operated satisfactorily.

Ferret Reconnalssance System

Ferret Reconnaissance System poyloads are being devel-
oped in a minimum number of configurations, The designa-
tion ond purpose of eoch configuration is as follows:

F-1 — R&D Test Payloads

F-2 — Digital General Coverage Pay-
loads

F-3 — Specific Mission Payloads

a .
CONF La irik
Pa] WDLPM4 188



STORED
TMACE

SCALE
1:528, 000

‘WDLPM-4 188

36 INCH LENS

LENS - 36" FOCAL LENGTH
ALTITUDE-300 STATUTE MILES

EXPOSURE~- 1/100 SEC. AT F/2.8

SCALE - 1:60,000
1 MILE

CONTACT PRINT
ILLUSTRATING
SCALE OF IMAGE
100 MILES

17 MILES

9 X ENLARGEMENT 300 X ENLARGEMENT

Figure 7. Simulated photography from satellite vebicle.



Figure 1,
Drawi of AGENA "BY
Propalilon sgitem

The AGENA vehicle was originally designed by the
Air Force for use as the basic satellite vehicle for
the Advanced Military Reconnaissance Satellite Pro-
grom (Weapon System 1171). The vehicle was
designed to be boosted by an ATLAS ICBM and
basic dimensions were derived from this booster
selection. The type of trajectory possible using the
ATLAS booster, coupled with the stringent eccen-
tricity requirements of these programs, led to
selection of a satellite guidance system suited to
accomplishing orbital injection in a horizontal atti-
tude. This led to the development for the AGENA of
an optical Inerticl system for vehicle guidance and
Gas jets for orbital attitude control. The Bell Aircraft
LR81-Ba-3 engine (Bell Hustier engine developed for
8-58 aircraft) was chosen for AGENA Propulsion due
to its advanced state of development. The YLR81-
Ba-5 version of this engine was developed to provide
increased performance through the use of unsym-
metrical di-methyi hydrazine {UDMH) fuel instead of
JP-4. Accelerated flight testing of the AGENA ve-
hicle and its subsystems became possible when the
DISCOVERER program was created, bosed on the

WDLPM-4 183

low cost and early availability of the THOR IRBM,
‘and a study which indicated that the AGENA could
be flight tested successfully in low altitude orbits,

Progress in the design of payloads for the MIDAS
Program created an urgent need for the attainment
of higher altitude orbits. As a result a modification
Program was initialed to develop the AGENA “BY
configuration. This work was authorized by ARPA
Order No. 48 and continued under ARPA Order No.
96. On 17 November 1959 Program responsibility
was assigned to the Air Force by the Secretary of
Defense. The AGENA “B Configuration includes the
addition of a single restart and extended burn
capability to the ‘engine and propellant tankage of
twice the AGENA “A” Capacity. The engine (Bell
Aircraft Model 8081}, for use on this vehicle, is
officially designated USAF Model XLR81-Ba-7, A
subsequent version, XLRB1-Ba-9 {Bell Aircraft Model
8096), with @ nozzle expansion ratio Of approxi-
mately 45:1 will further increase Performance capa-



CONFIDENTIAL

MONTHLY PROGRESS— ABLE Projects
ABLE-4 THOR
Flight Schedule

© Welding repairs made necessary by @ leak in
the oxidizer tank outlet caused a delay of several
days in the ABLE-4 THOR launch date.

Note

The ABLE-4 THOR was launched success-
fully from the Atlantic Missile Range on
11 March, All three stages performed
excellently and the payload continued to
be tracked along the programmed trajec-
tory. A detailed report of the launch and
orbital performance of the payload will
be given in next month's report,

Pre-launch Operations

@ An extensive testing and checkout Program for
ABLE-4 THOR systems and components was culmi-
nated during February. Forty environmental tests, 25
acceptance fests, 2 type tests, and 12 RED tests were
completed during the month. Three of the most sig-
nificant tests performed were:

@. Indications of high level third stage engine
vibrations in the 600 csp range required testing to
determine the effective mass measurements in this
Grea. Test results were coordinated with NASA,
resulting in revision of the specifications. On a basis
of the revised specifications, the back-up flight pay-
load was redefined as the type test payload and
vibration tested fen times in each direction at high
level, and at very high level at 600 cps.

b. The flight payload also was subjected to vibra-
tion acceptance testing and to additional tempera-
ture tests.

¢. Altitude soak tests were performed on the
Payload at a simulated 600,000 feet and at high

+ temperatures. Only minor difficulties were encoun-
fered in two of the experiments, both of which were
traced to test instrumentation and were corrected.

@ Following high altitude soak testing, the flight
payload was shipped to AMR on 24 February. All
recommendations of the Space Technology Labora-
tories Design Review Board had been accomplished
and all tests required by NASA had been completed.

Payload Batteries

© Protonged overcharge tests of the payload bat-
teries revealed a pronounced unbalance of charge

CONFIDENTIAL

current into the two parallel sections of the battery
pack. The following approaches were used in on
effort to alleviate this problem:

4. The balance of current on charge was added
to the battery selection criteria.

b. Two tempercture sensing thermal switches were
added to each battery section and set to
operate at 120 degrees F. The two switches
for each section are wired in parallel and
connected in series with the two switches for
the opposite section. :

Extensive environmental, vibration and reliability
tests were conducted on the thermal switches, using
the standard solar cell source and battery set-up.

© The battery manufacturer indicated at this time
that the maximum operating temperature for obtain-
ing satisfactory battery life should be revised from
150 down to 107 degrees F. As a result, afforts
were initiated by STL to improve the battery cooling
methods. Control of battery temperatures within the
revised limits was subsequently realized by replacing
the standard melamime sheet battery enclosure with
@ thin (2-3 mils} coating of epoxy, baked for 17
hours at 95 degrees F.

@ Additional battery tests were then performed,
simulating frequency of load applications and
charge rate under estimated flight conditions. The
test was conducted for 6 hours under the following
conditions:

a. Melamime sheet battery enclosure.

b. No thermal switches included.

¢. Approximately 80 degrees F. ambient and
vacuum temperatures.
Under these conditions, the severity of the test was
increased considerably. Test results were highly
satisfactory,

Payload Experiments

@ An electronic system was constructed for supply-
ing various types of currents to the Helmholtz coils
used for magnetic checkout of payloads at AMR
(Atlantic Missile Range}. This equipment will be used
for mapping the magnetic induction vector on ABLE-4
THOR payload. Recalibration of ABLE-4 THOR mag-
netometer has been completed. The tests were also
run to measure the ripple pickup by the search coil
from the power mains, These tests were necessitated
because of the rewiring of the payload harness.

WDLPM-4 188


Project MERCURY represents ihe transitional threshold viding ten ATLAS boosters modified in accordance.

between this nation’s cumulative achievements in space with program objectives and pilot safety factors, and
research and the beginning of actual space travel by (b) determination of trajectories and the launching
man. The primary program objective Is to place a and control of vehicles through injection into orbit.
manned satellite into orbit about the earth, and to The division of responsibilities for this Program is given
effect a controlled re-entry ond successful recovery in Table 1. Specific details of AFBMD support.are given
of the man and capsule (Figure 1). Unmanned ICBM in Table 2.

trajectory and near-orbital flights, ond unmanned Major contractors Participating in the AFBMD portion
orbiting flights will be used to verify the effectiveness of this program include: Space Technology Labora-
ond reliability of an extensive research program prior tories, systems engineering and technical direction;
to manned orbital flights (Figure 2). The program will Convair-Astronautics, modified ATLAS boosters;
be conducted over a period of nearly two years. The GE/Burroughs, ATLAS guidance equipment; and Rock-
initial R&D flight test was Sccomplished successfully etdyne, engines. All of these companies also provide
in September 1959. The total Program accomplish- special studies and engineering efforts peculiar to

ment is under the direction of NASA, The primary meeting Project MERCURY requirements.
responsibility of AFBMD to date consists of: (a) pro- .

——_ __
LEGEND —— ©
RESPONSIBILITY: = ‘,
AFBMD ==acs :

NASA =——

ESCAPE ! y
ROCKET | ,
7
! 7
CAPSULE / y
f v () BALLISTIC TRAJECTORY
BALLISTIC TRAJECTORY
aTLas iy TO SIMULATE REENTRY CONDITIONS
BOOSTER if NEAR ORBITAL INSERTION, IMMEDIATE
i REORIENTATION, DECELERATION, REENTRY
I) ATLANTIC RECOVERY

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

Figure 2, Plight test trajectories for Project MERCURY, defning specific objectives, ‘Trajectory C
serena ius bath ofthe final (manned) light. The point wt which APBMD aad NASA responsiblicg

i) —CONFIDENTFIAL WDLPMA4 108


Figure 6, Cuteway view of ABLE-4 THOR flight test vebicle (lefs), Drawing (right) shows ertist's
conception of ABLE-4 THOR payload in orb ches tn sae

WDLPMA 168 CONFIDENTIAL



A. THOR—SM-75 / AGENA “A"

Flight History

8 THOR—DM.21 7 AGENA “8” C. THOR—DM-21 / AGENA "B”

MB-3 Block 1 XIR81-Ba-7

‘MB-3 Block 2 XLR81-Ba-9

oe Ne. bow Plight Doje Remarks
o 1019 160 21 Janvar AGENA deswoned
” THOR reforbis set woh boo ase ow inde

1 1022 163 28 Feb 1959 “ne heciget eve sceutly Telemetry received for 314 seconds hier
it 1018 170 13 April ale 17th orbit

‘ea satnrrcon ga cette te 25 71 ok
am 1020 174 34 3b, ed orbital boost |. Failed

“ne io athe bi hese of tee pia bots mcs engine
Vv 1023 179 25 June Same as DISCOVERER It.
v 1029 192 13 jeved” sale

August Fe cltecioes succesfully achieved except capsule recovery after
vi 1028 200 19 August Same as DISCOVERER V.
vil 1051 206 7 November ti ed fully, Lae Lack of 400-cyele power prevented st.”
vil 1050 212 20 ber Attained orbis tion ed ACENA

Novem! ie aay pany bled soli petreeee prevented eens
Xx 1052 218 4 id

February zor shes dows Peale pol vo ue rapa area id wot retract.
x 1054 223. 19 February THOR destroyed ai T plus 56 see. by Range Safety Officer.

WY:
WDLPM-4 188 AS


APROX. . AGENA SECOND STAGE
MAK yin saet-7 engine
ATLAS-D BOOSTER
Thrust (Ibs at sea level)
‘Main Engines (2) 309,000
Sustainer Engine (1) 57,000
Vernier Engines (2) 2,000
ot Total at lift-off 368,000
“ Figare 2, Plight vebicle for STEER and TACKLE
Y phase.
{
i)
'
{
2
STEER objectives include: COMMUNICATIONS SUBSYSTEM—The three ele-

a. 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 north-
ern latitudes.

b. Development of engineering concepts and
equipment, and furnish test support data for DECREE
phase.

c. Investigation of the effects of vacuum and

tadiation 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 modi-
fied AGENA TMB’ 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 apoges 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 altitude control system will
then orient the Payload antennas toward the earth
‘ond solar cell paddles toward the sun to permit
communications system operation.

ments 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 maxi-
mum reliability in the possible Presence of interfering
signals. Initial test antennas will be 40- and 60-foot
Parabolic types. These will be replaced with hard-
ened or semi-hardened antennas later in the program
to provide compatibility with SAC hardened con-
trol centers. The receiver will use the same antenna.
The ground antennes 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 modu-
lation {spread spectrum) will be tested during this
phase.

Since the function of the satellite repeater is the real-
time relay of messages, separate reception and
transmission frequencies (10 to 15 megacycies apart)
Gre to be used. Satellite transmissions will be ot a
40-watt r-f level, using FM modulation and a 10-
decibel gain antenna.

When properly oriented toward the earth, the satel-
lite antenna coverage is approximately the some as
the angle subtended by the earth at a 5,600 nautical-
mile altitude. Coded contro! circuitry in the satellite.



WDLPM-4 188

—SNFDENTAL

DISTRIBUTION
Headquarters, United States Air Force %
Headquarters, Air Research and Development Command : 10
Strategic Air Command
Air Force Cambridge Research Center
Air Force Flight Test Center

1
1
1

Rome Air Development Command 1
Air Force Missile Defense Command 1
Wright Air Development Division 1
Air Force Special Weapons Center 1

’ Air University 4
‘Amold Engineering Development Center 2
Air Proving Ground Center 6
Air Defense Command a]
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
6555th Test Wing (Development) 2
Air Force Ballistic Missile Division (ARDC) 7
Ballistic Missiles Center (AMC) YW

Assistant CINCSAC (SAC MIKE} 3
Aeronautical Chart and Information Center 1
Rend Corporation

San Bernardino Air Materiel Area 3
6594 Test Wing (Soteltite) 1
1002 Insp. Gen, Group 1



The ARPA COURIER Program consists .of two flight
vehicles to be launched from the. Atlantic. Missile
Range. The program objective Is to test delayed
fepecter communications between a satellite and
ground stations. The program also will be used to
determine the operating characteristics and capa-
bilities 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 vehicle, paylood 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 re-
sponsible 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 contro! is exercised by a gyro platform
and a programmer. The second stage is controlled by
@ gyro used to govern engine gimballing during pow-
ered 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 «60-inch sphere, containing radio repect-
ers, 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 burn-
out, 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 nau-
tical 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 exchang-
ing “real time” information when the satellite is within
line-of-sight of two ground stations. During these peri-
ods a ground station can relay messages direct to the
next ground station, through the satellite simplex
repeater equipment.

Ground Support Stotions—These stations will be
located et Camp Salinas, Puerto Rico; Torrejon Air
Force Base, Madrid, Spain; and Halemano, Hawail.
Station design and development is under contract to
International Telephone and Telegraph Corporation.

79.3"

‘CONTBENTIAL-

60'

"(7 |e [e [a]

|
—

cl

wu Yhucieme

SECOND STAGE—ABLE-STAR {AJ10-104)

‘Thrust ot altitude 9030 pounds
‘Specific impulse (vor) 270 seconds
Total impulse {min} 2.3 X 10° ths/sec
Burning time 294 seconds
Propetiont + Uquid

PIRST STAGE—THOR IRBM.

Thrust fs. Ly 151,500 pounds

‘Specific imputse (s. LL} 248 seconds
Specific Impulse {voc} 287 seconds
Summing time 158 seconds
Propeliont Liquid



‘WDLPM-4 188

MONTHLY PROGRESS— COURIER Program
Launch Schedule

©@ The launch of COURIER 1A has been delayed
from June to 15 July becouse of payload availabil-
ity. At the request of ARPA, the launch of COURIER

1B also has been delayed and is scheduled for 1 _

September. Modifications to AMR launch pad 17B
for COURIER vehicles is proceeding on schedule.

Technical Progress
COURIER 1A

© Delivery of the second stage propulsion system
(AJ10-104) to the Spoce Technology Laboratories
Hangar in Los Angeles is scheduled for 25 March.

COURIER 1B
®@ Manufacturing and delivery problems relating to

the Minneapolis-Honeywell gyro have been resoived.
“Stop Work’ orders issued to the contractor during
January have been cancelled.

© The initial release of all drawings and specifica-
tions té Aerojet-General Corporation and Space
Electronics Corporation was compieted during Febru-
ary. Fabrication of subsystems has been started by
these contractors.

© Space Technology Laboratories will provide
Aerojet-General Corporation with systems checkout
van #2 and associated hangar checkout equipment
for use during the Azusa hangar checkout of the
ABLE-STAR systems. A complete list of all special
test and test support equipment that will be provided
to Aerojet-General Corporation and Space Elec-
tronics Corporation by STL was issued during
February.



WDLPM4 188

CONFIDENTIAL

MONTHLY PROGRESS —Project 609A
Program Administration

© launch Schedule — The lounch date for the
first NASA SCOUT flight test is now scheduled for
“not eariler than June.” This date reflects the most
current of several delays from the original October
1959 launch date. Previous NASA delays of long
duration -have resulted in a lesser slippage of the
609A schedules. However, it is intended that the
Present faunch date of 11 May 1960 will be main-
tained for the first Project 609A flight test vehicle. An
investigation is being conducted of methods of
Preserving 609A schedules in the face of continuing
NASA slippages. .

©@ Funding — NASA has stated that additional
funds are required immediately to complete the work
currently under contract for Air Force Project 609A,

Force Delivery Ord, » The increased costs have
been particularty significant in the guidance and
Control areas. The funds requested by NASA would
have exceeded the funds available, not only for
FY 1960, but the total funds progrommed for the

of items on order have been teduced, The total pro-
grom has been reduced also, to nine flight tests,
including two less expensive unguided vehicles
instead of the Previously scheduled one,

609A Booster Support
Requested by Navy

@ the Department of the Navy (BuWeps) has
requested Alr Force Project 609A booster vehicle

tional test phases of the Navy TRANSIT Program,
AFBMD has furnished support capability information
to AFDAT. In accordance with the 18 September
1959 Secretary of Defense memorandum assigning
booster support responsibility for space programs to
the Alr Force, Hq USAF (AFDAT) has coordinated a
“Navy Plon for TRANSIT“ and committed the Air
Force to support the Program. The Air Force com-
mitment imposes a significant emphasis on the need
for maintaining Present Project 609A schedules
regardless of further NASA slippoges to make a reli-
able vehicle available in time for the TRANSIT
launches. The Navy is Gniicipating the imminent
transfer of this program from ARPA, Approximately
18 Atlantic Missile Range launches will be required,
using 609A vehicles to place 50- to 100-pound pay-
loads into 500-mile circular orbits. The first launch is
required about May 1961 with following launches on
@ one a month schedule.

Technical Progress

@ Six static test firings of the modified ABL X254
(third stage) engine have been conducted success-

Figure 3. Project 609A best sage vebicie (Aerojet Senior)
being mated with second stage (Thiokol Sergeant, improved).

M3


7

Figure 2. ABLE-3 third sage ond pepload with solar paddles
ins fully extended position,

The ABLE Projects Program consists of the develop-
ment and launching of three vehicles from the Atlan-
fic Missile Range. The program is being ‘conducted
by AFBMD under NASA direction. The equipment
and objectives of the three flights are as follows:

ABLE-3—This four stage flight vehicle was launched
from the Atlantic Missile Range on 7 August 1959.
The vehicle consisted of a THOR booster, a second
stage using the AJ10-101A rocket engine, a third
stage powered by the ABL-248 A3 engine, and a
fourth stage consisting of the payload and an injec:
tion rocket. In addition to carrying a highly sophisti-
cated payload, the ABLE-3 flight was used to
demonstrate the validity of the ABLE-4 vehicle and
component configurations. All phases of the launch-
ing were successful and the advanced scientific
observatory satellite was placed in an extremely
elliptical geocentric orbit about the earth. Trajectory

Figure 3. Drawing of extremely elliptical ABLE-3 orbis.

Ba CONFIDENTIAL

WDLPM-4 188


WDLPM-4 188

ADVANCED SYSTEMS STUDIES

1, The Advanced Systems Studies Division has sev-
‘eral 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, spoce has been divided into three broad
Greas; earth orbital, lunar, interplanetary. Studies
in the lunar and interplanetary area are being man-
aged and directed at AFBMD. There are two studies
in the Lunar area: SR 192 {U) Strategic Lunar Sys-
tem, and SR 183 (U) Lunar Observatory. There is one
study in the interplanetary area: SR 182 (U) Strategic
Interplanetary System.

2. 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 opera-
tions such as offensive, defensive, ond supporting
systems. This study was funded with $600,000 in
Fiscal Year 1959 and final reports from the con-
tractors are due at AFBMD 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 esti-
mated as equivalent to one million dollars.

* 3. An obvious military requirement in the lunar area

will be a surveillance and intelligence collection sys-
tem. 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 intelli
gence observatory on the moon from which the
entire earth ond its surrounding area can be kept
under continuous surveillance, All earth orbital sys-
tems 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 ii
mated that this study has the equivalent. of $1.5
millions being applied to it.

4. The interplanetary area is being studied under
SR 182 (U) Strategic interplanetary System. The
objective of this study is to determine the possible
military missions and the type of equipment neces-
sary 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 lim-
ited, 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. Con-
frectors' final reports are due at AFBMD in Feb-
ruary 1960.

Nl



el=l*T-TeP[o[s[ep~

WDLPM-4 188



CONFIDENTIAL

Able-Star _

This program will develop a versatile and effi-
cient upper stage for use with varied booster/vehicle
combinations. This stage will have basic design fea-
tures proven in the dependable AJ10-101 stage
used on the THOR/ABLE vehicles, Improvements
being made include: (a) increased propellant capac-
ity; (b) a multiple restart capability; and (c} a full-
time attitude contro! system fo operate during coast
Periods as well as powered flight, These improve-
ments will permit a two-stage THO2/ABLE-STAR vehi-
cle 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,
Hf needed,

StU NE NAL CEN TENE MRA OSI ARAN nate gato

Figure 1. Side view of ABLE-STAR sebicle mousted in handling dolly.

WDLPMA 168 CONFIDENTIAL El


billly"The AGENA "8" configuration has increased Paylocds may be initalied on the AGENA forward
the effectiveness of the payload weight/orbital alti- equipment rack, or distributed throughout the vehicle. -
tude relationship for this vehicle. This reduction of An ejectable recovery capsule has been developed
Payload restrictions has increased Greatly the poten- “for the nose section of the AGENA vehicle and flight
tialities of using the extremely reliable THOR as an tested in the DISCOVERER Program,

AGENA booster, In order to obtain maximum efft-

ciency from the use of ATLAS as an AGENA booster,

modifications fo the ATLAS are being made to reduce NOSE CONE

limitations imposed by trajectory requirements and
static toad capabilities,

NOSE CONE

32

FUEL AND
wae /—FUEL AND
OXIDIZER
SECTION bev
SECTION
SEPARATION
PLANE SEPARATION
PLANE
LIUM
SPHERES orients
“ULLAGE ROCKETS if aa ULLAGE ROCKETS
NGINE (UDMH) -- SPHE NGINE (UDMH)
RGENA TAT CNA TET
Tayo 8 Weight—Inert A Yeoo,
6,550 Impulse Propellants 13,100
Fuel (UDMH)
Oxidizer (IRFNA)

67 Pyrotechnics 100

oF “5
7,987 GROSS tbs.)
rating...
ae


: ENTER ELLIPTICAL ORBIT .
PERIGEE OF 7
COAST ELLIPSE. a
a
AGENA \ Y
‘st, aan

‘ ‘AGENA 2 ND. BURN
SEPARATION & ENTER

SHORT COAST

+ Figure 2, attitude reference. Also. governs velocity ma,
Laxnch-to-orbit trajectory for flights 3 and subsequent, Opti- direction by inertial guidance system monitor by horizon
mum ATLAS boost, guided by radio-inertial system. Scanner. Orbital attitude maintained by reaction wheel and
AGENA ascent (coat, barn, coast, second burn) provides Gas jets.

PROGRAM HISTORY

The MIDAS Program was included in Weapon
System 117L when WS 117L 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 en infrared reconnaissance system.
The MIDAS {Missile Defense Alarm System)
Program was directed by ARPA Order No.
38, dated 5 November 1958 until transferred
fo the Air Force on 17 November 1959,
Development activities will lead to the first of
@ ten flight R&D progrom in February 1960,
with a reliable operational system achievable
by 1962-1963,

Figure 3.

Propo MIDAS eT OO satellites

swo gonal planes at ‘Wms. abies;

erage of USSR with midimsm number of
c2 prs

distant in each
way Protas mecitane et



— 2 oo

SHIGNLS IDvdS

. .



. BOR tiem tw
rr ONES

serie version $f sowed presen

tation sytem (les) for dspley of
dite is ditpleyed om 4 TV onic

t

y Can
determine when an actual attach bas
been launched, The decision to
counterattack is made by the Presi-
dent, with all afected agencies react-
ing as preplanned.

The MIDAS system is designed to provide
continuous infrared reconnaissance of the
Soviet Union, Surveillance will be con-
ducted by eight satelite vehicles in occu.
rately positioned orbits (Figure 3). The area
under surveillance must be in line-of-sight
view of the scanning satellite. Mission capa-
bilities are shown in Figure 4. The system is
designed to accomplish instantaneous read-
out of acquired data by at least one of

SORBET
E a
ey ON cn

three strategically located readout stations,
The readout siations transmit the data
directly to the MIDAS Control Center where
it is processed, displayed, ond evaluated
(Figure 5.) if an attack is determined to
be underway, the intelligence is communi-

cated to a central Department of Defense _

Command Post for relay to the President
and all national Fetaliotory and defense
agencies.


SW3LSAS IDWS

sysfems

COMMUNICATIONS
SATELLITE

DISCOVERER
SAMOS
MIDAS



WDLPM-4 188

an
“Dont ENTIAL

MONTHLY PROGRESS — TRANSIT Program
TRANSIT 158

@ The second stage vehicle is scheduled for ship-
ment to AMR on 12 March following completion of
various assembly and checkout Procedures,

Launch Schedule

@ Launch date has been rescheduled from 5 April
to 12 April because of the following work require-
ments:

a. Replacement of helium bottles and engine
onti-rotation bolts found to be defective on the
second stage propulsion system.

Rework of the second stage nitrogen attitude

control system manifold to eliminate undesire-

able design characteristics.

© Increasing the size of the explosive bolt holes
(Stages 1/11).

d. Rerouting of an oxidizer jine.

@. Relocating the Stages 1/11 electrical con-
nectors,

f. Providing a mounting bracket for the Propul-
sion electrical sequence unit.

b,

Testing and Checkout

© Altitude tests have been completed which inves-
tigated the possibility of the payload surface becom-
ing contaminated from fuel vapors which might be
Present in the helium gas used to actuate the pay-
load spin and separation mechanism. Test results
indicate that no contamination problem exists.

©@ The Stages I/II structural tests were completed
during February.

© Second stage systems checkout van #3 was val-
idated at Los Angeles and is to be returned fo AMR.
The second stage launch control consoles also have

been validated and are being used in hangar Oper-
ational tests of the vehicle in Los Angeles. The con-
soles will be shipped to AMR for installation.

TRANSIT 2A
Launch Schedule

@ Launch date hes been delayed to 18 May
because of the delay in the TRANSIT 1B schedule.
This date permits a five week interval to be main.
tained between the two flights.

Technical Progress

© ‘The defective helium bottles and engine anti-
rotation bolts reported for TRANSIT 1B were also
encountered on the TRANSIT 2A second stage. These
units are being replaced.

© The second stage flight propulsion unit (AJ10-
104) was delivered during February. The equipment
compartment has been instalted and aligned and the
forward and aft electrical cables installed. Also, the
thrust chamber assembly and attitude control jets
have been checked and aligned.

TRANSIT 28

©@ Launch date has been set tentatively for October
or November, depending upon the availability of the
ground support equipment for the second stage
advanced guidance system.

@ = Manufacturing and delivery problems reported
last month for the Minneapolis-Honeywell gyro have
been resolved. “Stop Work" orders issued to the
contractor in January have been cancelled.

© The initial release of all drawings and specifi-
cations to Aerojet-General Corporation and Space
Electronics Corporation was completed during Feb-
tuary. Fabrication of subsystems has been started by
these contractors.

“ONT

“FS


S19310dd Ives

projects

AGENA
ABLE-STAR
MERCURY
609A

COURIER
TIROS


Al® FORCE BALLISTIC MISSILE DIVISION (ARDO)
UNITED STATES AIR FORCE
Alr Force Unit Post Office, Los Angeles 45, California

‘WOLPM-4 25 March 1960

ForEworp

Activities summarized in this report include the major space systems, projets
nd studies for which the Air Forea Baltic Mile Division is wholly or por.

. Progroms must be sufficiently flexible to permit continuous and effective
integration of rapidly occurring advances in the state-of-the-ort,

PLeibi

i

i

i

i

i Pos. RITLAND
‘Commander
i

L

L

l

L

L

i



CONFIDENTIAL

PROJECT MERCURY

———$——_

NASA Responsibility
Booster-Capsule integration
Capsule
Ground support
Recovery
Flight 1 Flights
roa *T
Laboratory Flight
Capsules
(McDonnell)
!
Re-entry
heating,
dynamics,
recovery

Unmanned non-orbitat 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 Ibs,

Orbit altitude 105-115 nm,

Automatic pilot safety equipment with

manval backup

AFEMD Responsibility

|

Booster
ATLAS “D"

Modified

INew guidance antennae

Thin skinned tanks

Shifted tank bulkheads

Light weight thrust shructure
Capsule separation mechanism
Interface electrical connections
LOX dome protection
Programmer Modified

Range safety destruct time delay
jAbori-sensing system —_-

Detects dangerous variations
Launch in tank pressures, fuel injector
Pressure, electrical power,
attitude and sustainer hydraulic
Pressure. Can initiate escape
sequence

Integrated Sountdown—including pilot insertion
ond 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.
Approximate launch azimuth:

la} Bollistic frajectory—108*

(b) Orbital trajectory—072°

Table 1. Outline of NASA and APBMD responsibilities in PROJECT. ‘MERCURY,

WDLPM-4 198 CONFIDENTIAL


A. THIRD STAGE-—X-248 (Allegany Ballistic Lab.) T

Thrust ot altitude 3190 pounds A
‘Specific impulse (vac) 250 seconds =
Total impulse 116,400 Ibs/see

Burning Time “97.5 seconds

Propellant Solid B

8. SECOND STAGE—AJ10-42 (Aerejet-General) _ |

Thrust at altitude 7700 pounds
Specific impulse (voc) 271 seconds
Total impulse {onin} 870,000 tbe/sec
Burning time 115 seconds
Propellant Liquid

C. FIRST STAGE—THOR RBM

Thrast (a) 151,500 pounds
Specific impule (6.1) 248 seconds
‘Specific impulse (vac) 287 seconds
Burning tee 158 sqconds
Propellant Liquid

ay)

Figure 1, TRANSIT IA three stage flight vebicie,

TRANSIT IA' launched from Aslantic Missile Range

The TRANSIT Program consists of the flight testing of
four vehicles to place 200-270-pound satellite pay-
loads into circular orbits of 400 to 500 nautical
miles. The program is designed to provide extremely
accurate, world-wide, all-weather navigational infor-
mation for use by aircraft, surface and subsurface
vessels, particularly in relation to POLARIS missile
firings. The ARPA Order for TRANSIT 1A was initiated
in September 1958 and amended in April 1959 to
odd TRANSIT 18, 2A and 28 flights. The program is
currently authorized by ARPA Order No. 97, which
cssigns AFBMD responsibility for providing the
booster vehicles, integrating payloads to the vehi-
des, ond flight operations from launch through
attainment of orbit, including communications to the
tracking and data handling facilities. Payload and
tracking responsibility has been assigned to the

Figure 2. Casvauey drawing of TRANSIT 1A payload (NAV 1). oni the of ond nance. Applied Physics Labora.

a a | m1



WDLPM-4 198

CONFIDENTIAL

MONTHLY PROGRESS—Tinos Program
Program Administration

@ The revised Powered flight trajectory was op-
Proved by NASA and a revised Detailed Test Objec-
tives document issued. All vehicle design changes
and modifications required by the new trajectory
have been completed and no delay in the launch
schedule is anticipated.

@ Trajectory and initial Impact Point data have
been submitted to range safety for approval and no
changes in range safety requirements or the Mill-
stone (Massachusetts) beacon antenna locations or
patterns are anticipated at this time.

Flight Schedule

©. A launch date of 31 March was established dur-
ing February, keyed to the following prelaunch
schedule:
First stage on stand
Second stage on stand .
Third stage on stand’.....

T-6 days test . 24 March
T-1 day test .. 30 March
Launch ... ee 31 March
Technical Progress
© System design and qualification tests were com-
pleted successfully during February.


40. SECOND
ASPECT
INDICATOR TIMER

BATTERIES INJECTION ROCKET

HEAT SINK
COMMAND BOX

DIGITAL 5W TRANSMITTER
DECODER

PROPORTIONAL
COUNTER
(CHICAGO)

STATIC
CONVERTER

106 MC
TRANSMITTER

WDLPM-4 188

TELEVISION
LOGIC CIRCUIT

COMMAND RECEIVER

SOLAR CELL .
CURRENT MONITOR

FLUX GATE
MAGNE TOME TER

THOR-ABLE I PAYLOAD

SUBCARRIER OSCILLATOR
AND
MULTIPLEXER "8'

BATTERIES

SUBCARRIER OSCILLATOR
MULTIPLEXER "A"

DIGITAL
TELEMETRY

(TOP VIEW)

Figure 4. ABLE-3 payload (top view).

and orbit were essentially os Predicted with devia-
tions in apogee and Perigee occurring on the more
than nominal side. The Payload was the most sophis-
cated to have been placed In orbit by this nation
ot the time and contained provisions for conducting
13 experiments in Space environment and Propaga-
Hion. A wealth of valuable dato was obtained from

sufficient power for franamitting signals due to solar
paddle damage suffered during initial paddle exten-
sion and the resultant unfavorable sun "look" angle.

ABLE-4 ATLAS—This vehicle differed from the
‘ABLE-3 only in that an ATLAS ICBM was used as the
first stage instead of a THOR (RBM. The unsuccessful
launch of the ABLE-4 ATLAS occurred on 26 Novem-
ber 1959. Structural breakup resulted in the third
stage and payload parting from the vehicle approx.

CONFIDENTIAL


WDLPM4 re CONFAD

discrepancy was indicated in the Pulse
width measurement circuits. The circuit design is
being studied in an effort to solve this problem.
Efforts to solve the intermittent time counter errors
encountered during systems testing of the third F-1

payload are Progressing - Desensitizing

the counter stages appears to be the most feasible -

solution, A breadboard of the desensitized time
counter has been installed in an Fel service test
medel payload and has been Operated satisfactorily
for 48 cycles of life testing (equip to approximately
three days of orbital operation). The use of line
filters is being studied as an additional effort. Sepa-
ration tests of the vehicle nose cone were completed
satisfactorily during January. Separation tests simu-
lating vehicle-p attachments were completed
satisfactorily during February.

@ F-2 and F-3 Payloads — in accerdance with
the fam reorientation reported in raph
8.1, ‘he concephe ‘and basic characteristics for the
new F-2 and F-3 payloads were defined in an LMSD
Technical Letter Report for January. Work statements

sign and modification of some of the payload com-
ponents affected by the change {i.e. payload struc-
ture and antenna assemblies) have been initiated.

foot cable (center of photo).

Ground Support Equipment -

© Delivery of the F-1 data conversion equipment
to the Sctellite Test Center is scheduled for 25
March, Negotiations ore Underway for the changes
to the F-2 and F-3 ground data handling equipment
resulting from program reorientation.

Program Communications and Control
Equipment

© Design of the exit VHF antenna for the satellite
vehicle has been refined, using a honeycomb dielec-
tric to support the cavity. A weight reduction of 60
Percent was realized and laboratory tests indicate
satisfactory performance,

© Systems and acceptance tests are being con-
ducted on the UHF Ground equipment for the Van-
denberg AFB tracking and data acquisition station,

Ground Support Program
@ Ground Handling and Service Equipment —
Equipment for Point Arguello launch pad #1 has

been delivered and is scheduled to be completely
installed and checked out by the middle of May.

@ Launch Control Equipment — Manufacturing
of launch control systems equipment for Point
Arguello launch pad 2 is 80 percent complete. The
equipment for launch pad 1 was shipped to Vanden-
berg AFB on 18 February.

Figure 8. F-1 payload separation test set.
ap. Satisfactory payload separation wat
demonstrated in tests conducted byLMSD
iu February. Payload-vebicle attachments
were simulated in test set-up. The zero
‘sresity condition of the payload in orbit
twas effected by suspending it trom « 60-

B7


PROZCT MERCURY
——l
AFBMD Responsibility
in support of
PROJECT MERCURY
NASA HS-36
includes:
ST
Design, engineering studies Launch support
Equipment modification Trajectory dota
Hardware fabrication Missile allocation
Flight scheduling
Provide ten (10) Modify boosters for NASA Prelim. Launch, control and define trajectories
ATLAS boosters, inary research and manned orbital of booster-capsule vehicle up fo, and
flight and safety objectives, including, injection into orbit.

Table 2. AFBMD responsibilities in support of PROJECT MERCURY.

MONTHLY PROGRESS—Project MERCURY
Flight Test Progress

@ The second launch in ATLAS/MERCURY flight
test program, designated MA-1, is scheduled for late
May or early June from Atlantic Missile Range launch
complex 14. MA-1 launch plans have been revised
by NASA and were Presented to \AFBMD during
February. The revised flight plan includes the fol-
lowing characteristics:

Booster — ATLAS 50D

Capsule — McDonnell structural shell with
external shingles and adapter. Internal
‘equipment will be furnished by NASA,

Trajectory — A near orbital trajectory will
be used, similar to that shown on Figure 2,
item B. Capsule abort from an orbital launch
will be simulated at “pproximately 20,000
feet per second.

Abort Sensing System — The abort sensing
system will be carried on the ATLAS vehicle
and will be open loop tested on this flight.
The escape system, designed as port of the

-CONFIDENTIAL

basic capsule configuration, will not be car-
ried on this flight, However, a stub escape
tower will be corried for ballast effect. Sepa-
ration of the capsule from the ATLAS booster
will be effected by posigrade rockets at.
tached to the capsule,

Primary Test Objective — To demonstrate
the structural integrity of the capsule air-
frame, Particularly the shingles on the coni-
cdl afterbody {see Figure 3). the shingles
are designed to provide structural support
and capsule insulation functions,

Ground Support Progress

© A meeting was held between the AFD ATLAS
guidance contractors {General Electric Co, and
Burroughs) and the NASA ground communications
and data link contractors (IBM, Western Electric Co,
Bendix, and Milgo} on 11-12 February. The Purpose

"of the meeting was to reach agreement on the Bur-

roughs-iBM interconnection specification for provid-
ing NASA with ATLAS/MERCURY guidance data.
This data will be used by the NASA Contro! Center
and computing facilities ot Beltsville, Maryland,

WDLPM-4 188


eee
PROSECT 609A
Hyper-Environment Test System

ABL X248 AG
PROGRAM DESCRIPTION —The Hyper-Environment
p02 JETS Test Program (607A) is divided into RED ond Opere-
tional Phases. The R&D phase will be used to develop
TT ‘ond flight test vehicles capable of carrying 50 to
ABL X254 A2. 1,000 pound payloads to altitudes of 200 to 7,000
Hg 0» VETS niles, The Operational phase will use this stondard-
2 zed vehicle to permit the economical performance
on — 30" Of flight test experiments In wupport of scientific
research and advanced mallitary space system
Programs,
+ Sconomy—Relabillty Versatility —in this order
Of empbaals are the three significant guides to pro-
‘Gram occomplichment, ECONOMY Is being achieved
_~ THIOKOL
SERGEANT
20' (IMPROVED)
XM33E5
(ALT. NOZZLE)
Hy 0, JETS
ABL X248A6—_
- 3" :
ABL X254 A2:
AEROVET
SENIOR TM~
33KS 120,000
enone
30' RO FIN THIOKOL NY
SERGEANT Figo Thee srs
of Project 609A oebicle
8. NOZZLE cays sinew
. JET VANES AND
40 AERO FINS

WDLPM-4 188 CONFIDENTIAL M1


CONDITIONS;

2000 wm. elttede

8 vehicles in polor orbit
Faglaed
North Atlontie
North Pacific

b Obsiing satellites de
} Orbiting satellites detect in
edb

emit
Data selemetered instantaneously to MIDAS

ber of missiles launched and launch location,
. direction of travel and burning characteris.
‘ ” hes, above shows  fercentege of detec.
: tion probabilities over USSR.

WDLPM-4 188

ee sereneeneseaeeenreree

TECHNICAL HISTORY

The MIDAS infrared reconnaissance Payload
will be engineered to use a standard booster-
satellite launch vehicle configuration. This
configuration consists of a “D! Series ATLAS
missile as the first stage, and the AGENA
vehicle, powered by « Bell-Aircratt rocket en-
gine, as the second, orbiting stage (Figure 1).
Refinements fo the AGENA vehicle will be
made as a result of the DISCOVERER flight
test program. The first flight article infrared
Payload has been assembled and installed
on Gn AGENA vehicle, and checkout opera-

Cage Se

tions 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 -communica-
fions package, also has been designed, fab-
ricated, and tested. The total Payload weight
is approximately 1,000 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 attitude polar orbit (see Figure
2). Only the first two R&D flight tests will use
the single capacity AGENA. vehicle.