ATLAS
BOOSTER

15,100
74,900
172,00
262,300
356,000 e
$2,100
86
no
AQENA.
“an —
170, 1400
6,590 13,100
CA 100
797 14,800
YURS1-Be-S MURBIo7
18900 15900
a a7
120 FY)
No Yes
PROGRAM HISTORY
Figure 1. The SAMOS Program was included
rit comet of SAMOS tlie, pq 147. en rented fo oe
Line feavin o complete flight
vehicle (righ (right) and detailed view of Advanced Research Projects Agency
basic A NA upper stage (left). early in 1958. ARPA separated
WS 1171 into the DISCOVERER,
SAMOS and MIDAS programs with
the SAMOS objectives based on a
NOSE CONE viswal and ferret reconnaissance
system. On 17 November 1959
responsibility for this program was
_ transferred from ARPA to the Air
Force by the Secretary of Defense.
PROGRAM MISSION
The primary mission of the SAMOS advanced
env reconnaissance system is to provide continuous
TANK visual, electronic (and other) surveillance of the
section USSR ond its allied nations. Efforts include devel-
opment of hardware to permit:
a. Determination of characteristics of enemy
electronic emissions,
PLANE b. Verification of known targets, detection of
unknown targets,
¢. Location and evaluation of defenses.
d. Evaluation of military and industrial
LUM ‘strength,
‘SPHERES

@. Assessment of high-yield weapons damage.
f. Reconnoitering of troop movements.

9. Location of naval forces throughout the
world.

Bl



The ABLE Projects Program consists of the develop-
ment and launching of three vehicles from the Atlan-
fic Missile Range. The program ts being conducted

by AFBMD under NASA direction. The equipment
and objectives of the three flights are cs 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 AJIO-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-
Piya i ABLE 3 shied stage and payload with solar paddles ing were successful ond the advanced scientific
observatory satellite was placed in an extremely
elliptical geocentric orbit about the earth. Trajectory

ORBITAL PERIOD |:

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

Ea CONFIDENTIAL WDLPMA 200



WDLPM-4 200

Figure 3. Artist's concept of propored UHP communications satellite,



Studies

S3IGNLS 3D vas



Figure 6. Cuteway view of ABLE-4 THOR Might test vebicle (lef), Drawing (right) shows arsist's
conception of ABLE-4 THOR Payload in orbit about the son,

‘CONFIDENTIAL.



:Figare 3. Typical DISCOVERER trajector; (above)
from bie at J cadenbers APB re orbit. Spe)

satellite orbital path around 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

1 with technical manage-

ment assigned to AFBMD. On 14 November 1959, s WW

contractor for the program is Lockheed Missile and

Spoce Division. The DISCOVERER Program will pro- Primary objectives indude:

vide: (a) space research in support of the (a) Flight test of the sateltite vehicle airframe, pro-

military reconnaissance satellite systems, Programs, Putsion, guidance and contro! systems, Guailiary

(b) test of the ground communications and tracking anid snl, ond telemetry, tracking ond com-

network for these 2 ht testi mond equipment,

Of the AGENA second stage vennate, “WaMt HHng Attaining satelite stabilization in orbit,
. {cl Obtaining satelite internal thermal environment

data,

vous wo mer



WDLPM-4 200

CONFIDENTIAL

MONTHLY PROGRESS—COURIER Program
Program Administration

® A major revision of the Program Requirements
document for projects TRANSIT and COURIER has
been published.

Technical Progress
COURIER 1A

@ The AJ10-104 propulsion system which was to
be delivered on 15 March will not be delivered until
8 April. This late delivery will not change the hard-
ware availability for launch.

@ The late propulsion system delivery created a
problem in the availability of checkout vans. Van 2
which is used in Los Angeles will be in use by
Aerojet-General Corporation on COURIER 1B. This
will necessitate van 3 being shipped from AMR to
Los Angeles to permit checkout and assembly of the
second stoge vehicle. Van 3 would then be shipped
back to AMR with the vehicle.

@ * Some delay in the calculation of the final flight
trajectory was coused by the late propulsion system
delivery. This trajectory is needed for establishing
the proper rates for the control system and times for
the programmer.

@ it was determined that inadequate aerodynamic
heating protection for the guidance antennas had
been provided. The application of an ablating coat-
ing will solve this problem.

© Aerodynamic heating studies indicate that some
protection will have to be afforded the payload
because of nose fairing. radiation. The COURIER
payload contractor has agreed that aluminum foil
can be used to eliminate the additional heating
caused by the nose fairing radiation. The foil must
Rot cover the inner surface of the nose fairing lower
than the payload equator.

COURIER 18

@ Space Technology Laboratories is continuing to
release drawing changes to Aerojet-General Cor-
poration and Space Electronics Corporation as
chenges occur on the COURIER 1A configuration. No
schedule changes due to vehicle design are antici-
pated at this time.

©@ Aerojet-Generai Corporation and Space Elec-
tronics Corporation are continuing fabrication of
components and subsystems based on a scheduled
launch date of 1 September.


+. . Holsted onto tower platform
(left) . . . being lowered toward
second stage (center and right).

Hoisted onto gantry (left)... being mated with third stage (center)
+++ Installed on third stage (right).

“CONFIDENTIAL EIS


Payload in High Altitude
Chamber Tests

Foliowing extensive rework of the
paytoad early this year it was
subjected to severe operational and
environmental testing in the high
altitude simulation chamber shown.

WDLPM- 200 CONFIDENTIAL Eis


CONFIDENTFIAL-

feet per second faster than Necessary to escape the
earth's gravity. First acquisition of the vehicle bya
Ground station in the ABLE tracking network was
made at approximately 0815 EST by the Jodrell Bank
Experimental Station at Manchester, England.

Parameter Nominal Actual
Inertial burnout

velocity 36,597 ft/sec 36,404 ft/sec
Inertial velocity

angle ° 75.80° 78.28°
Inertial velocity

azimuth 93.04° 93.02°
Latitude 28.105° 28.25°
Longitude {from

Vernal Equinox) 297.6° 297.9°

Radial distance from
the earth's center 22,275,000 ft 22,336,000 ft

Table 1. BURNOUT PARAMETERS

Flight Test Vehicle

© First Stoge—Stand THOR intermediate range
ballistic missile, with the inertial guidance system
removed and a modified control system. Guidance
was accomplished by roll and pitch programmers,
The Rocketdyne engine used RU-1 fuel and LOX.

@ Second Stage—ABLE vehicle. Aerojet-General
Corporation (AJ10-10tA) propulsion system using a
hypergolic combination of inhibited white fuming
nitric acid oxidizer and unsymmetrical dimethyt
hydrazine fuel. The second stage incorporated an
STL Advanced Guidance System utilizing a ground
based Burroughs Corporation Mod I guidance
‘computer,

©@ Third Stage — Allegany Ballistics Laboratory
248-A4 solid propellant rocket, with a burning time
fo propellant depletion of 38 seconds. Spin stabilized
by six spin rockets installed on the second stage.

Satellite Flight

@ The PIONEER V satellite is presently on a 152
day journey of 46,400,000 miles to reach its closest
point to the sun on its elliptical solar orbit. The
closest point to the sun (perihelion) will be about
70,000,000 miles. This point is only a few ‘million
miles away from an intersection with the orbit of
Venus. Distance from the earth will increase to
approximately 186,000,000 miles when the satellite
reaches a point on the far side of the sun (estimated
date: December 1962). Its orbit will then swing the
satellite back to within 50,000,000 miles of the earth
sometime in 1963. Its galactic velocity in solar orbit



‘WDLPM-4 200

CONFIDENTIAL

ated approximately 15 percent of the time; the 150-
watt transmitter can be operated approximately 2
Percent. of the time. The nominal power available
from the solar cell conversion system is about 15
watts when the payload is near the earth and will be
about 24 watts at its perihelion.

@ The batteries operate at 18 volts. A series of
static converters provides a variety of voltage levels.
An undervoltoge control automatically removes the
transmitter toad from the batteries in case battery
discharge threatens to disable the receiver and thus
Prevent the satellite from being commanded from
the ground. in addition, the undervoltage contro!
@ssures the conservation of battery lifetime by pre-
venting deep battery discharges. Battery voltage is
ranging between 16 to 19 volts and temperature is
varying between 80 to 100°F.

Experiments

@ The payload aboard the PIONEER V satellite
consists of experiments designed to measure three
types of phenomena in space: radiation, magnetic
fields, and micrometeorite density.

Radiation Experiments

@ Designed to determine the relative abundance
of the different species of charged particles and to
indicate their energy distribution. Although each of
these experiments is independent, the results cannot
be evaluated separately. Although all three sensors
measure the total flux of particles, each has a differ-
ent detection threshold. Total fux is measured by
‘each sensor above its particular threshold. The com-
bined sesults provide an intensity versus energy
spectrum.

@ lon Chamber and Geiger-Mueller Tube (Uni-
versity of Minnesota)—The ionization chamber is of
the integrating type used extensively in the University
of Minnesota balloon flights. Together, these two
instruments provide information giving the incident
Particle flux ond mean ionization per particle. The
minimum energy particles which can enter the detec-
tors are about 1.25 mev electrons or 25 mev Protons.

@ Proportional Counter Telescope (University of
Chicagol—This experiment contains a bundle of
seven small proportional counters; arranged with a
central counter surrounded by two sets of three
counters. It is designed to provide single incidence
and triple coincidence measurements to obtain a
measure of both primary cosmic-rays and the low-
energy background. Results, to date, correlate with
count rates obtained from EXPLORER VI in its apogee

areas. In addition, two occasions of sudden decrease
in cosmic-ray intensity have been observed which
correspond to similar decreases at ground monitor-
ing stations and one probable decrease which does
not correspond. There is evidence of an increase in
“soft radiation several hours after the decrease in
cosmic-ray intensity. The experiment continued to
function excellentty as this reportwas being prepared.

Magnetic Fleld Experiment

@ Search-Coil Magnetometer—This unit is
mounted near the periphery of the satellite to meas-
ure the component of the magnetic field perpendic-
vlor to the spin axis. By making use of the spin of
the vehicle {knowing the orientation of the axis of
the spin) the magnitude of the magnetic field through
which the satellite is moving can be deduced. A
Phase comparator which measures the phase relation-
ship between the output of a photodiode sun scanner
(aspect indicator) and the output of the magnetom-
eter permits accurate knowledge of the direction of
the field in the spin plane. Performance of this
experiment continues to be excellent at the time the
report is being prepared. Early results essentially
confirm those obtained from PIONEER f and EX-
PLORER VI. This experiment has defined a point
which seems to be the actual “edge” of interplane-
tary space. Beyond this point, approximately 64,000
miles from earth, no evidence of a magnetic field
appecrs to exist.

Micrometeorite Experiment (Air Force Cambridge
Research Center}

© The micrometeorite mémentum spectrometer
determines the times of micrometeorite impact and
separates the momenta into two groups of different
energies, Equipment consists of a diaphragm on the
payload shell, a microphone mounted beneath the
diaphragm, an amplifier, and a pulse height ana-
lyzer. Comparison of the total count of microme-
teorites encountered with the two-value level of
energy measured will permit an evaluation of the
density of particles encountered and also, estimates
of their energy {velocity or size). Experiment perform-
‘ance has been excellent but data has not been ana-
lyzed sufficiently to draw any scientific conclusions.

Astronomical Unit

@ In addition to the scientific results made possible

by the specific experimental equipment carried on
PIONEER Y, the precise tong-range tracking data will

CONFIDENTIAL

E23



‘WDLPM-4 200

Search Coil Magnetometer

Search Coil
Magnetometer Logic

PIONEER V
PAYLOAD

(bottom view)

Phase Comparitor *

150-watt Transmitter S-watt Transmitter

“CONFIDENTIAL El


WDLPM-4 200

CONFBEN THA

MONTHLY PROGRESS—SAMOS Program
Second Stage Vehicles

@ Systems checkout of the AGENA vehicle for the
first SAMOS flight is in progress at the System Test
Area (formerly Modification and Checkout Center).
Shortage of o satisfactory flight article UHF narrow
band transmitter will couse a delay in shipment of
this unit to Sante Crux Test Base for static hot firing
tests. This delay will not cause a rescheduling of the
+2 September 1961 launch date. The AGENA vehicle
for the second SAMOS flight was delivered to the
System Test Area for systems testing late in March.
Assembly of the major structural components of the
AGENA vehicle for the third flight was accomplished
on 9 March. The vehicie is now in the final assembly
phase of manufacturing. The focilities checkout vehi-
cle for the AGENA “A” (dual payload configuration)
was shipped to Vandenberg AFB early in March.
Design of the AGENA “B” vehicle for the fourth
SAMOS flight is proceeding on schedule. This vehicle
will carry an E-2 camera and will be the first single
Payload flight in the SAMOS program. Design
requirements ore being established for the AGENA
vehicle programmed to carry the E-5 (recoverable
visual) payload. Seven such flights are scheduled at
this time.

Visual Reconnaissance Systems

@ Visual Reconnaissance System payloads are
being developed in a minimum number of configura-
tions to attain readout and recovery mission objec-



Figure 1. ABLE-3 flight test vebicte beis launched from Atlantic Missile . Dimensional drewin,
(left) of fowe-stage. ‘ABLES vebicle. being Revre *

El


WDLPM-4 200

“CONFIDENTIAL

‘The ARPA COURIER Program consists of two flight
vehicles to be launched from the Atlantic Missile
Range. The program objective is to test delayed
repeater communications between a satellite and
ground stations. The program also will be used to
determine the operating characteristics and 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, payload integration, launch,
injection of payload into orbit, and verification of
orbital parameters at injection. The Army Signal
Research and Development Laboratory will design,
develop and fabricate the payload, and will be re-
sponsible for world-wide ground station requirements,
Primary payload contractor is Phileo Corporation,
Vehicle Description—The two-stage COURIER vehicle
Consists of a THOR booster, an ABLE-STAR (AJ10-104)
second stage and 2 500 pound COURIER payload.
Booster flight control is exercised by a gyro platform
and a programmer. The second stage is controlled by
Q 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 nitire acid
and unsymmetrical dimethyl hydrazine. The engine
will have a restart capability.-The 500 pound COURIER
payload is a 60-inch sphere, containing radio repeat.
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-
aut, 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 nav-
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 Stations—These stations will be
located at Camp Salinas, Puerto Rico; Torrejon Air
Force Bose, Madrid, Spain; and Halemano, Hawail,
Station design and development is under contract to
(nternational Telephone and Telegraph Corporation.

[= |o =| [¢]o [al

79.3"
H
60"
i]
'
‘
'
‘
J
T
SECOND STAGE—ABLE-STAR (AJ10-104)
* Thrust of altivde (0030 pounds
Specific impelse vor) 278 seconds
Total impulse {min} 2.3 X 10° ths/sec
Burning time 294 seconds
Propeliant Liquid
FIRST STAGE—THOR inam
Theust fs. 1) 151,500 pounds
‘Specific impolse {4.1 248 seconds
Specific Impulse (voc) - 287 seconds
Quming time 138 seconds
Propetiont Uquia

CONFIDENTFIAL—

Gl



arly In 1958, the decision to accelerate the national space effort was made effec:
‘ely postible only because of the availabilty of the THOR IRBM, THOR No, 127
wos diverted from the R&D flight test program for use as the ABLE-1 space probe
fint stage. With top national priority aetigned to the space research effort, THOR
Ne, 163.was used fo booet the DISCOVERER I into orbit on 28 February 1959. Since
then, the THOR hes become a reliable operational IRBM and highly reliable aleo ox
@ booster for space vehicles. During 1959 all THOR boosted space flights achieved
completely successful first stoge performance. THOR performance has been increased
through weight reduction. modifications and use of RU-1 (instead of RP-1) fuel, In
April 1960 a modified THOR, designated DM-21, wil be available, incorporating a
shortened guidance compartment and additional weight reduction changes. In July
1960 THOR thrust will be {i :
Increased to 167,000
Pounds through installation
Gf the MB-3-Block Il engine.
The first DM-21 vehicle wil!
be used to boost DISCOV-

Prime. contrecter:

Coavale .

Engine menvtacterer:
=e,
Height A tect
Olemeter tees
Weight 261,206 pounds

The ATLAS ICBM, Providing over twice the thrust of the THOR, will be used os
the fit stage booster forthe three Advanced Miltary Satelite Programs and for
Project Mercury man-in-space. The first ATLAS boosted space flight was launched
from the Atlantic Missile Range on 18 December 1958, Designated Project Score,
this vehicle (ATLAS 10B) successfully placed a communications Payload satellite
into orbit around the earth. In November 1959 the ABLE~4 space probe did not
attain orbit; however, the ATLAS first stage performance wos entirely successful.
The first flight test vehicle in Project Mercury was launched on 7 September with
alt objectives easenticlly achieved and excellent ATLAS booster performance
realized. Future flights will use modified ATLAS series “D" missiles to carry
increased payload weights. Project .
Mercury boosters are being modified
also to include abort-sensing and
other pilot safety features. The success
Of the first three ATLAS boosted space
flights, plus the. increased perform-
ance and reliability being demon- -
strated in the ATLAS R&D flight test

Program, lend -confidence in this

booster as a means of realizing

‘WDLPM-4 200



WDLPM-4 200



“CONFIDENTIAL

The TIROS Program consists of one fight from the
Atlantic Missile Range early in 1960. Primary objec-
tyes Includes {a) To determine the feasibility of using
‘an earth satellite to mecsure,.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-
de, 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 date.

VEHICLE DESCRIPTION

The three-stage TIROS vehicle (Figure 1) consists of
a 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.

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

IY Thrust at altitude 3150 pounds
Specific impulse (vac) 250 seconds
Totel impulse 116,400 Ibe/see
Burning Time . 37.5 seconds
Propellant Solld
SECOND STAGE—AJ10-42 (Acrojet-General)
Thrust ot altitude 7700 pounds
Specific Impulse (voc) 271 seconds
Total impulse (min) 870,000 Iba/sec
Burning time 115 seconde
Propeltant Liquid
FIRST STAGE—THOR IRBM
Threst (1.1) 151,500 pounde
Specific impulse (1) 248 seconds
‘Specific impulse (vac) 267 seconds
Burning time 158 seconds
Propellant Uguid

WDLPM-4 200

FLIGHT DESCRIPTION

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 380 nautical mile circular orbit
having an inclination angle of
48.67 degrees. Orbital life js ex-
Pected 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
9260 solar cells to recharge the
batteries, The solar celts, installed
in the top and cylindrical side walls
of the satellite will furnish an 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 resolution capabilities and coverage pat-
terns 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 « ground station, pictures may be
taken on command and transmitted directly to earth.
‘When the satellite is beyond radio communication
range, camera operation is controlled by a clock
and programming circuits and the images recorded
‘on magnetic tape for readout during the next pass
over a ground station. Two beacon transmitters are
installed on the bottom side of the satellite to facili-
tate trocking.

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 mcs and com-
mand transmission on 140 mes. Required modifica-
tions to the TL 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 attaining an
early buildup of experienced personnel,

CONFIDENTIAL

[w [o [+ [=|]



‘trouble within the satellite vehicle. At T-2 days the
satellite was removed from the stand and the launch
was rescheduled to 1 April. The repairs were made,
the satellite was reinstalled, and the checkout Pro-
ceeded with no further trou! + The launch count
down was conducted without a hold until the stort
of the terminal count (7-35 minutes). At T minus 2
minutes a hold was initiated because of diffculties
encountered with the “C” band beacon. The count
‘Was recycled first to T-8 minutes and then to T-15

froze. These problems were corrected and the launch
‘was accomplished on schedule.

@ All three stages of the TIROS vehicle Operated
within the programmed limits, with the exception of
the “C” band beacon. However, the operation of
this beacon did not affect the orbit obtained. For
the nominal and actual Powered Trajectory Flight
Plan refer to Table 1.

@ The Fort Monmouth, New Jersey, and Kaena
Point, Hawaii stations are tracking, commanding and
receiving data from the TIROS satellite as planned.

© This fight has Proved the feasibility of using an
earth satellite to provide information for more accur-

ate long-range weather forecasting. The SAMOS
and MIDAS programs wilt benefit because of this
successful flight. Unofficial information received indi-
cates the successes accomplished by this satellite.

a. High quality pictures taken by the wide angle

camera are being received on both the direct

‘and delayed read-out modes.

High quality pictures token by the narrow angle

Camera are being received on the direct read-

out mode only, The delayed read-out mode is

not operating property.

¢. The solar cell recharging rate is higher than
anticipated.

d. All tracking, command, and data read-out
transmitters and receivers are operating satis-
factorily.

b,

@. Some malfunction in the system is Preventing
Proper sun scanner (attitude indicator) opera-
tion,

© Data received indicates that the Payload will be
©petational throughout its Programmed useful life
of five months. The satellite will remain in orbit for
several years.

woe ray Et

Vertical lift-off ° 0
BTL guidance begins Stage! closed-loop steering 90 90
End Stage! pitch program 130 129.55
MECO—Stage I main engine cutoff 158.5 160.80
Stage #! separation and ignition; begin Stage it pitch program 162.3 164.70
BTL guidance begins Stage il closed-loop steering 172.3 175.80
BIL guidance ends Stage it closed-loop steering 253.9 250.90
BTL discrete ends Stage I pitch Program 255.5 252.80
BTL discrete spins up third stage and payload—120 pm 268.6 265.70
Stage if cutoft—SECO; begin coast Period 270.6 267.70
Stage lit separation; start Stage Ill coast timer 272.1 269,20
Fire Stage til rocket 662.89 664.0
Injection into orbit 700.31 699.0

Table 1. Powered Trajectory Flight Pian

H4

WDLPM-4 200


C8

mitting @ substantial increase in the amount of data
that can be transmitted over a single transmission
link because of its extremely fast sampling rate
140,000 samples per second). Data quality also
is better than conventional telemetry commutation

@ A study hes been made of a defection monitor
for radar antennas. The monitor would Permit a
reduction in the highly rigid design requirements for
radar mounts by Providing a deflection correction
system for less rigid mounts. This would permit the
use of lighter metals and simpler designs, and result
in lower antenna mount costs.

@ Each radar beacon for flights 2 and subsequent
will be adjusted to respond to an individual interro-
gation or command signal transmitted by the Verlort
radar. This procedure will permit individual commu-

nication with each MIDAS satellite when more than
one satellite is within range of a tracking station
at the same time. The Verlort radar, designed to
operate simuttaneously with six different beacon
addresses, will require no modification.

Facilities

@ The MIDAS Program addition to the Vandenberg
AFB data processing and acquisition building will be
completed on 15 April,

@ During March, an Air Force site survey team held
conferences with local authorities in Salisbury,
Norther Rhodesia, Africa, regarding possible loca-
tions for MIDAS Program tracking sites, The survey
of Northem Rhodesia will take @pproximately 4
weeks. This station will provide data on the AGENA
“B" restart portions on flights 3 and subsequent.



‘ “+ | ORBITAL
iL COAST PERIOO—"T" Boost OReiy .
\ wari. ——
i PLATFORM | PLATFORM
u WMERTIAL ¢ COMPUTER wos Ste ADY- gr,
we AEORIENTATION ATE ORB ~~
L paaction J6TS jevRo Con ENTIAL.
REMTON SETS | ae moons” ATTITUDE
\ REACT,
| aun amin 6 Cactiygy ON WHE:
irs ee Tive) el
4 — |
i cee
a ADORE OF ascent se, ae
: STATUTE MILES)
é. be
_ Figure 2. Typical SAMOS trajectory, lawnch- to-orbis.
L - Ferret Reconnaissance...

‘|

Figure 4. The Perret reconnaissance system will gather. data from electronic emissions over unfriendly territory.

~ Ba SECRET- -WDLPM 200



© The first Model 1604 computer will be delivered
to the Satellite Test Center on 1 April. Checkout is
scheduled for completion by 15 April. The second

computer will be delivered to Vandenberg AFB on
15 June.

Facilities

@ Vandenberg Air Force Base — Missile
Assembly Building final acceptance inspection was
completed 16 March. The technical support and
laboratory buildings construction contract has been
awarded. Additional LMSD requirements for the
data acquisition and processing building have
resulted in a slippage of the beneficial occupancy
date which could affect the launch schedule. Con-
struction is being expedited and installation of
equipment concurrently with building construction
being considered.

@ Point Arguello—tounch pad 1 first acceptance
inspection wos completed 18 March, and the facility
is ready for occupancy. Technical support building
plans ond specifications were released to the con-

struction: agency on 18 March, with a beneficial
occupancy scheduled for late July.

© Satellite Test Center—Construction of incre-

ment 2 is scheduled for completion during mid-June.

Design of a modification to this facility that will
accommodate a revised equipment configuration has
been initiated. Completion of construction is sched-
uled for October.

© Offutt Alr Force Base—Construction of the
interim data processing facility will be advertised
in April. Construction of the interim facility is sched-
uled for September completion. Design of the com-
plete SOC/DPF is being held in abeyance until a
firm SAMOS program justifying this facility is estab-
lished.

@' New Boston, New Hampshire, Tracking
and Data’ Acquisition Station—Construction of
technical facilities for this station is complete and
ready for installation of equipment. Equipment deliv-
eries .will not be made during 1960 for planned
interim activation of this station in late 1960.

@ Ottumwa, iowa, Tracking and Data Acqui-
sition Statlon—Plans and specifications for the
technical facilities for this station are complete and
teady for advertising. Design of support facilities for
this station is being held in abeyance until a firm
SAMOS program requiring the Ottumwa station is
established.

rst

“WDLPM-4 200


Figure 6. SAMOS concept, showing reception of commands and transmission of data
. between satellite and ground station; and Subsystem functions (schematic),

For economical testing of components a dual-
capability visual and ferret payload will be used dur-
ing the early development phase. On later flights
only a visual or ferret system payload will be carried.
These payloads will be housed in the AGENA vehicle
(Figure 1), *

Dato collected by the visual payloads will be elec-
tronically transmitted in the readout system ond
retrieved in the recovery system. Ferret data will be
transmitted electronically. These systems are com-
posed of the AGENA vehicle, ATLAS booster, launch
facilities, tracking facilities, and a communications
‘and data processing network. The recovery system
will also include a re-entry capsule and a recovery
force.

CONCEPT

ATLAS Series D missiles launched from VAF8 will”

boost the AGENA vehicle into polar orbits. Injection
into near-circular orbits (Figure 2) will be accom.
plished by the AGENA vehicle rocket engine. A self-
contained guidance system using a horizon reference
scanner will provide altitude stabilization. As the
satellite travels in an orbit essentially fixed in space
the earth rotates.inside the orbit (Figure 3). Each

successive orbit is displaced laterally @pproximately -

23% degrees at the equator, permitting one vehicle
to observe the entire earth in a time period depend.
ent upon the width of the area under surveillance.
Early versions will have a useful life of approximately

ten days. The readout systems will hove a usetul life
of four months with a design objective in certain
configurations of one year; recovery systems will
have a useful life of fifteen to thirty days.

TECHNICAL DESCRIPTION

Visual Program—Payload camera, film processor
and electronics readout equipment are being devel-

* oped by Eastman Kodak Co. Cameras having a
36-inch focal length are being used, The payload .

equipment includes automatic film Processing, film
transport and take-up, electronic readout and tem-
perature controls. The recoverable system will retain

both the exposed film and the 66-inch local length

Ferret Program—Ferret Payloads are being devel-
oped on @ progressively more advanced basis from
R&D (F-1) to advanced systems (F-4). The F-2 ail-
digital, general coverage payload will use super-
heterodyne scanning receivers in conjunction with
directional antennas, an ‘analog to digital converter
and tope recorders {for storage). A Programmer will
be used to contro! read-in over areas of interest and
readout over tracking stations. The F-3 payload will
use similar receivers with stop-scan capability and
controllable antennas added. Recording of the actual
signal intercepted {rather then the digital represen-
tation) will be possible with'a bandwidth up to é6mc.
A complex programmer will permit satellite search
of a given area or frequency range.



The Communications Satellite Program will Investi-
gate the feasibility of using synchronously spaced
satellites as instantaneous repeaters for radio com-
munications. Under ARPA Order No. 54, as
amended, AFBMD is responsible for the design,
development, and. fight testing of the complete
system, including launch, satellite tracking and con-
trol, and necessary support facilities and ground
equipment. The Army Signal Research and Develop-
ment Laboratory has been delegated development
management responsibility for the microwave com-
munications subsystem as directed by ARPA Order
54,

The Communications Satellite Program is currently
being conducted in accordance with amendment 5,
{dated 11 April 1960) to ARPA Order No. 54. Under
this amendment the previous method of accomplish-
ing the program objectives in three progressively

Figure 1. Proposed satellite with jettisonable
fairing mounted on AGENA second stage.

JETTISONABLE
NOSE FAIRING
PAYLOAD MAXIMUM
va a ENVELOPE
“yt TM
Cf] )
SSL
3 | —SOLAR PADDLE
if \ }
\ i
NJ LY 7
fa “s
a
(as
a4 {_— AGENA STAGE
aa PTS
ene

Dl



WDLPM-4 200

successfully photographed by Baker-Nunn cameras
of the Smithsonian Astrophysical Observatory, pro-
viding the most precise satellite tracking data ever
obtained. Accuracy of the TRANSIT beacon and DIS-
COVERER verlort rodar stations can now be closely
determined by checking against this Optical data.
Both frequencies of the doppler beacon operated
and were tracked. Photos of the light were made by
the Cadiz, Spain camera station,

© Although capsule separation and retro rocket
ignition took place, the capsule did not descend in
the recovery area. Verification of the ejection for
re-entry sequence was indicated by telemetry,
although there was considerable noise at this time
and valves are somewhat difficult to determine, The
capsule telemetry transmitter was tracked by the
Hewailan station and the recovery aircraft received
brief signals. Intensive efforts are underway to deter-
mine what occurred and take corrective action prior
to the Jaunch of DISCOVERER XII.

© This flight is of great significance to the SAMOS
and MIDAS programs. The components that are crit-
ical for success of SAMOS and MIDAS functioned
extremely well. The modifications made in the past,
based on results of previous DISCOVERER flights,
were proven to be very effective. The flight accom-
plished the outstanding feat of achieving near-

perfect performance of all components except the
de-orbit function. Since an advanced version of this
satellite vehicle will be used for both SAMOS and

Figure 6. AGENA "B vebicle being loaded
WeND oh fon tbe DISCOVERER (DIS.

program a
COVERER XVI).

Pi 7. Hydraulic control aimbaling the XLR-
Bibed ‘engine, This fuck poones Lon cmbalng the XLR-
DISCOVERER fight XVII.

MIDAS programs, the development and refinement
of the complex satellite vehicle and subsystems at
this point in time lends confidence to the success of
these programs.

Techaleal Progress
Second Stage Vehicles

Differences in DISCOVERER vehicle configurations
are defined. on page A-5.

@ All AGENA “A” vehicles (for use on DISCOV-
ERER flights X! through XV) have been delivered to
Vandenberg AFB.

@ The first AGENA “8” vehicle (XiR-81Ba-7
engine) was delivered to the Santo Cruz Test Base on
1 March. The vehicle has been installed in test stond
2 and preparations are being made for hot firing
tests. The other three AGENA "B" vehicles using this
‘engine model ore'in the LMSD Systems Test Area
(formerly Modification and Checkout Center). Sys-
tems checks of the three vehicles are 90, 65 and 60
Percent complete.

© The test firing program for the XLR-81Ba-9
engine was completed during March at the Santa
Cruz Test Base. This engine, installed in an AGENA
“B" vehicle, will be used on DISCOVERER flights XVII
and subsequent. The test program consisted of nine

A7



‘CONFIDENTIAL—

by long range Planning and maximum integration mate maximum of 400 miles with 150 pound pay-
other programs. Use of the basic four-stage, loads,

solid propellant, SCOUT vehicle, developed by NASA AR -
and modified to achieve m 609A Program Management ‘An ‘abbreviated devel-
Progra objectives, pment plan, covering the R&D Phase only, was

will effect. an economy in vehicle development,
‘approved on 9 January 1959, Funds in the amount

Necessary modifications Include provisions for sta-
bilizing the fourth stage without spin ond use of the of $9,077,000 were made cyoilable for this abbre-
vehicl viated portion of the program only, A letter was issued

le in less than the full four-stage configuration.
aasigning management Fesponsibility to AFMBD, with

Close integration with the current ballistic missile
Program will effect on economy by Permitting tests emphasis on integrating the Program with the scien-
and experiments to be conducted on tegularly sched- tific and military research experiments conducted on
uled ballistic missile test fights whenever possible.” {aavlarly scheduled bolfistic missile fight test (iggy.
without delaying schedules, Economy in the opera- back Program). In dune 1959, Aeronutronic Division
tional phase will be exercised by the use Of this low. of the Ford Motor Company was chosen through nor.
cost vehicle as a standard flight test platform to mal competitive bidding as the Payload, Test, and
Perform scientific and military experimental research Systems Integration Contractor, Arrangements have
in support of all Air Force facilities, RELIABILITY wilt been made for the procurement of vehicle compo-
be obtained by a twelve vehicle R & D flight test nents and associated support equipment, modified
Program, at least four flights of the basic SCOUT, to meet Program 6094 requirements, through NASA,
and. maximum use of knowledge goined in Prior Air tather than through the SCOUT Program contractors,
Force ballistic missile flight testing, VERSATILITY will Atlantic Missile Range facilities consisting of launch
be achieved by designing a vehicle capable of being complex 18 will be made available to the Air Force
readily adapted fo a wide range of Payload varia- for this program. A Project 609A division has been
tions, and capable of being flown in several configu. established within the 6555th Test Wing (Develop-
rations of four stages or less. This VERSATILITY resuits ment) at AMR to supply Air Force technicians to par-
in the following flight capabilities: (a) vertical probes ticipate in the assembly, checkout and launch opera-
nevig a wide variance of payload welght/atttude Hons of the R&D phase under the direction of the
combinations; (b} boost-glide trajectories; (c) ballistic Payload and Test Contractor: An all-military opera-

ies; {d) downward boosted, high-
ini trajectories; { Ronal capability will be developed from within this

speed re-entry Profiles, and (e) full orbit to app

group,



CONFIDENTIAL |

MONTHLY PROGRESS—nRos Program
Flight Test Progress

@ The TIROS satellite was launched from stand.
I7A, Atlantic Missile Range, at 0640 EST on 1 April,
This launch again Proved the reliability of the THOR-
ABLE combination. Because of the Payload restraint

of obtaining a specific angular relationship with the

sun at injection the launch had fo be made within

@ one hour time span or be Postponed to the fol-
lowing day. The TIROS satellite was launched
Gpproximately three minutes before the end of that
Period,

@ All prelaunch checks progressed smoothly until
T minus 5 days. On 26 March, during checkout the
* payload de-spin squibs were fired. A faulty squib
fester was believed to have been the cause, but
further investigations showed it to be caused by

TIROS LAUNCH

TIROS leaving stand 17A, Atlantic Missile Range, on
1 “April for its successful orbit. This list of orbital
Parameters shows the nominal and actual valves

achieved.
Apogee 381.5N.M. 498.8 N.M.
Perigee 379.8 N.M. 378.3 N.M,
Period 98.67 min. 99.17 min,

Latitude ot injection 42.336°N 42.277°'N
Longitude at injection 56.90" W 57.045° W

Inclination angle 48.330° 48.359"
Payload spin rate 12 rpm 10 rpm
fin orbit)

THIRD STAGE—PAYLOAD

This picture shows the Joad mounted on the third 33

tebicle, The tor cells gree, manated on tb ‘perbbery Sf

the payload. Three of the four antennas can be seen Projecting
* down from the payload.

WDLPM- 200 CONFIDENTIAL Hs


WDLPM-4 200

CONFIDENTIAL

the reception at Manchester is teduced to one bit per
second. This reception rate is expected to occur when
the satellite is approximately 10,000,000 miles from
earth (about 15 May). When this condition occurs,
the 150-watt transmitter will be fumed on and is
expected to provide acceptable communications up
fo a range of 50,000,000 miles,

© For communications to the earth, the transmitter
accepts an I-F signal from the satellite receiver,
translates in frequency, and amplifies it either to a

5- or 150-watt level. When the 150-watt transmitter
is operating, the 5-watt transmitter acts as a driver.
In the. process, the signal (378 mc frequency) is
modulated with a 1024-cps subcarrier containing the
time-multiplexed pulse-code-modulated output of the
Telebit System. Biphase modulation is used to impress
the telemetry output on the subcarrier. The 150-watt
transmitter weighs about six pounds.

@ Performance capabilities for the ground stations,
computed prior to launch, are as follows:

Days Ronge Vehicle Information
Alter tN. Mi) Tranemitter Rote
Lawech —{Millfiens?— Power (watts! Station Uhits/ seed Remarks
5 7 5 Atlantic Missile Range 1 Maximum range
Singapore 1 for Singapore
Hewaii 8 and Atlantic
Manchester 64 Missile Range
14 2 5 Hawaii 1 Maximum range for
Manchester 8 5-watt transmission
to Hawaii
72 15 150 Hawaii 1 . Maximum range
Manchester 8 for Hawai
142 60 150 Manchester Jd Maximum range

PIONEER V Satellite Vehicle

~CONFIDENTIAL-

E21



SD scorn

eee a
THOR AB suzs | Boas, | ATLAS Seis
7 !
Weight—dry 7,74 6510 | Weight—wer 15,100 AS
Fuel 30,500 33,695 Fuel 74,900 i
Oxidizer 68,300 68,190 Oxidizer 172,300 i
TOTAL WEIGHT 106,546 108,395 | TOTAL WeicHT 262,300
Thrustbs,, $.L, 152,000 167,000 J Thrusttbs., $1, s
Spec. Imp.-sec, 246.42 248.3 Boost 356,000 fig
Burn Time—sec. 163.59 148.0 Sustainer 82,100 [PN
Spec. Imp.-sec, G
Boost 26
Sustainer 310
NOTES A G E N A D "AH ge Ss
E
© Payload weight not included. Engine Mode! YUR81-Bo-5 | X1R81-8o.7® XLRB1-Ba-9® c
Does include controls, guid.
ance, APU and residual pro- | ©Weight—inen 1,155 1,370 1,400 °
pellants, Impulse propellants 6,550 13,100 13,100 x
@ Does not include THOR Pyrotechnics 67 108 108
oan (2 ee, or ATLAS' @TOTAL WEIGHT 7,772 14,578 14,608 s
; Separation Weight 77H 14,552 14,582 T
@ Single restart copability. , A
Dual bur Thrust-ibs., vac. 15,000 15,000 15,000
® 8 operation, Spec. Imp.-sec., vac. 277 277 G
© Allegany Baliistic Laboratory. Burn Time—sec. 120 2400 E
; Ar AJ10-104
AEROJET- GENERAL 10-42 toto |ABLE-STAR °A BL 24 8
Weight—wat 1,247.1 847.9 1,297 59.5
Fuel 875.1 869.0 2,247 O 455.5
Oxidizer 2,499.6 2,461.0 6,227 (solid)
TOTAL WEIGHT 4,621.8 41779 9771 D 515
Bumout Weight 1,308.6 944.1 1,419 50.5
Thrustbs., vac. 7,670 7,720 7,900 . 3,100
Spec. tmp.-sec., voc . 267 267 278 250.5

Program Vehicle s Combinations

DISCOVERER {1 thru 16)...

DISCOVERER {16 thr 21),
DISCOVERER (21 thru 29)
COMM. SATELLITE

MIDAS (1 and 2).....,

SAMOS (1 thru 3)...
SAMOS (4 ond subs).

MIDAS (3 and subs}... .



useful (primary power) lifetime will be 24 days. A
launch azimuth of 107 degrees will be used. Launch
must take place between 10 a.m. and 3 p.m. EST.

@ Analysis of data obtained from the first MIDAS
flight has resulted in the following modifications
being made to the second MIDAS flight test vehicle:
G, Destruct system circultry modified.
b. Separation sequence modified.
¢. Instrumentation added in the ATLAS adapter
section to provide ftight information on com-
ponents associated with separation.
d. Transmission of adapter information changed
from AGENA to ATLAS telemetry.

@ in response to a Hq USAF request to evaluate
methods of conducting additional MIDAS program
flight tests between the presently scheduled second
and third MIDAS fights, AFBMD has proposed the
use of two DISCOVERER vehicles. MIDAS payloads,
using the backup infrared scanner units from the
fiest two MIDAS: flight vehicies, would be modified
fo permit installation in the DISCOVERER (THOR/
AGENA “B”) vehicles. This diversion of vehicles

Bi 7. Fall-scale solar auxiliary hep, be
lin foliation 1 “Uefe) aad fully te
zero gravity condi the artay ie orbis was
Gad oe support “der eh of photo).

~~

7



Flight Time Inertial Speed Downrange Altitude
Point (seconds) Comments (ft/sec) Distance (n.mJ (namJd
2A . 1.8 2A 1B 2A A 2-A
1B 2.3 Transit vehicles 28 28 28
10 10 End of vertical rise 1,346 1,346 0 0 0.077 0.077
2 167 167. First stage burnout 13,11 12,929 752 797 M2 48.3
End of second stoge
3 442 448 first burning period 24,539 24376 785.6 778.0 200.1 203.0
4 lap aa Festortuecond stage oo ugg aap 405324000 5000 5000
5 1,504 1,462 _ Injection into orbit 24,258 24,259 44163 4,130.0 500.0 500.0

FLIGHT TRAJECTORY —TRANsIr 18, 2A and 28

ene

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 lonosphere. 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 satellite at any point in the orbit
should be known. Using. known orbital positions,

WDLPM-4 200

Ss

ships and aircraft can then use satellite signals to
make analogous 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 Aifantic 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
as the closest sites possible to the orbit injection
point.

F3



MONTHLY PROGRESS—TRANSIT Program
Flight Test Program

© Avery successful TRANSIT 1B launch was accom-
plished from Stand 178, Atlantic Missile Range, at
0702 EST on 13 April. The THOR ABLE-STAR boosted
satellite was the seventeenth United States satellite
fo be placed into orbit.

@ The countdown proceeded smoothly except for
One two-hour hold caused by some minor vehicle
problems and satellite radio interference. The vehicle
problems were corrected and the cause of inter-
ference was found to be the NOTS infrared experi-
ment transmitter. The experiment would be terminated
after seven days in orbit and this would alleviate the
problem.

@ Both stages of the TRANSIT vehicle performed
within the programmed limits, The ABLE-STAR second
stage (on its first filght test} fired, then shut off and
coasted, then on command fired again. This was the
first time in the free world a rocket engine had been
restarted in space. See Table 1 for nominal and
actual orbital parameters,

Parameter Nominal Actual
Apogee 500 .N. Mi. 409 N, Mi.
Perigee 500 N. Mi. 175 N. Mi.
Period 103 Min. 95 Min.
Inclination Angle 49.95" 51.3°

Table 1, ORBITAL PARAMETERS

@ Tracking stations in the United States, Newfound-
land, and Germany are tracking and receiving data
from TRANSIT 1B,

© Experiments contained in TRANSIT 1B will help
to determine the feasibility of using a satellite as a
navigational aid and to obtain very accurate geo-
detic measurements. The success of TRANSIT 1B
Paves the way for future navigation satellites which
may very well obsolete century-old, weather-depend-
‘ent, celestial navigation, and the more recently
developed LORAN {long range aid to navigation).

TRANSIT 1B
Zi TRANSIT Bat ter Lito from ‘shad

FS



WDLPM-4 200

TRANSIT 1B .
Program Administration

@ The Dato Acquisition and Support Pian for
TRANSIT 1B, a major revision of the Program Require-
ments document for Projects TRANSIT and COURIER,
and the revised Detailed Test Objectives document
for TRANSIT IB, has been published,

Technical Progress

@ = The helium bottle and ‘engine anti-rotation bolt
fetrofit was accomplished prior to shipment in addi-
tion to the correction of the items listed in the
February report.

@ it was determined that adequate aerodynamic
heating protection for the guidance antennas had not
been provided. Treating the antennas with an ablat-
ing coating eliminated this problem.

Ground Support Equipment

@ The downrange telemetery and tracking station
underwent final system tests, including checkout of
the tracking function in conjunction with a manned
aircraft. The downrange station arrived at Erding,
Germany, on 21 March and has been erected. Call-
bration using a helicopter is scheduled for 1 April.

TRANSIT 2A

Program Administration

@ The Detailed Test Objectives document for
TRANSIT 2A and the Program Requirements docu-
ment for Projects TRANSIT and COURIER has been
published.

Launch Schedule

@ The assembly and checkout of the second stage
is proceeding on schedule based on a hardware
availability for launch of 18 May.

Technical Progress

© The rework of the helium bottles and the engine
anti-rotation bolts was accomplished.

@ = The aerodynamic heating problem reported for
TRANSIT 18 was also encountered on the TRANSIT
2A. These units will have an abiating coating applied.

© The thrust chamber assembly which was deliv-
ered with the AJ10-104 propulsion system was of
minimum performance. As a result or trajectory
studies using this performance the success of the
mission was in doubt. To correct this problem the
original thrust chamber assembly was replaced with
‘one of higher than nominal performance.

@ The second stage range safety, guidance and
control system checks have been completed. The
accelerometer ond Programmer have been checked
out on the vehicle. The propulsion contro! system
compatibility sequence test has aiso been completed.

Ground Support Facilities

© During March a site selection team completed
arrangements for locating a mobile downrange
telemetry and tracking van at Punta Arenas, Chile.
Arrangements have been completed for the local
telephone company to supply o teletype circuit to
the station.

TRANSIT 28

Launch Schedule

@ Because of interference problems on use of
ground facilities the launching of TRANSIT 2B was
rescheduled to early November.

Technical Progress

©@ Space Technology Laboratories is continuing to
release drawing changes to Aerojet-General Corpo-
tation and Space Electronics Corporation as changes
occur on the TRANSIT 2A configuration. No schedule
changes due to vehicle design are anticipated at this
time,



Figure 4.

emi
Data telemetered instantaneously to

Control Center via far north readout stations,
Decoded data is approximately the num-
ber of missiles launched and launch location,
direction of travel and burning characteris.

tics, above shows ye of detec-
tion Probables over USSR. ?

iting satellites detect infrared radiations
itted by Soviet ICBM’s in powered fight.

TECHNICAL HISTORY

The MIDAS infrared reconnaissance payload
is engineered to use a standard launch
vehicle configuration. This consists of a “D"’
Series ATLAS missile as the first stage and the
AGENA vehicle, powered by a Bell Aircraft

rocket engine as the second, orbiting stage .

{Figure 1). The total payload weight is
approximately 1,000 pounds.

The first two of the ten R&D flights will use
the AGENA “A” vehicle which is programmed

to place the payload in a circular 261 nauti- *

cal mile orbit. Subsequent flights will utilize
the ATLAS/AGENA “B” configuration which
will be programmed to place the payload in
a circular 2,000 nautical mile orbit.



‘WDLPM-4 200

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-
fime attitude control system to operate during coast
Periods as well as Powered flight. These improve-
ments will permit a two-stage THOR/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,
if needed,

Figure 1. Side view of ABLESTAR oebicle monnted in handling dolly

CONFIDENTIAL

Kl


ABLE Vehicle

The ABLE upper-stage vahicle has been flight tested suc-
cessfully as the second stage on THOR re-entry vehicle

ABLE Projects and TRANSIT 1A. The vehicle uses
AJ10-42 or AJTO-101 Propulsion systems (improved
versions of systems used originally on the Vanguard
Program), guidance systems, and electronic and instru-
mentation equipment. The ABLE vehicles are guided
during second stage engine burning. Vehicles using the

Contractor:

Aarolet-Generol Corp.

Height 18 feet 7 taches
Diameter 4 feet 8 inches
‘Weight

AN10-42 4622 pounds
4310-101 4178 pounds
Fuel

Unsymmetrical Dimethy! Hydrazine

Development of the Allegany Ballistics
laboratory X-248 engine for the Van-
guard Program was ‘accelerated when it
was selected as the third stage for Project
ABLE-1. The unit represented the most
advanced solid propellant engine of its
size available at the time. Since the
engine had not been qualification of
flight tested, test firings were conducted
in a vacuum chamber simulating approxi-
mately 100,000 feet altitude. Design
modifications involving the igniter,
nozzle, and internal insulation were
found to be required, The modified
engine performed with complete satis-
faction on the successful flight of ABLE-1

and subsequently on ABLE-3 and ABLE-4
THOR.

~CONFIBENTHAL-

AJ10-101 system are ‘spun with the third stage ond pay-
load prior to second stage engine burnout to provide spin
stabilization of the unguided third stage and payload.
On flight vehictes using the AJ10.42 Proputsion system,
only the third stage and payload Gre spun prior to second
stage separation by a spin table bearing system located
at the second to third stage séparation plone. Only minor
differences exist between the two Propulsion systems.

Onidizer
Inhibited White Fuming Nitric Add

Svidence
As0-42

Rodle-Inertlol (srl)
AlI0-101

Adtranced Guld. Syst, (sTt)
Compater @errovghe J-1)

Used as second stage for:
AN10-42 — TRANSIT 1A, Tiros,
AJ10-101 — ABLE 3 ond 4

ABL 248
Vehicle

Contractor:

‘Allegany Ballistic Laboratory
Height 4 feet 10 inches
Diometer 1 foot 6 Inches
Weight 515 pounds
Post Solid

Used as third stage on:
ABLE 3 and 4

TRANSIT 1A, TIROS


Although originally designed as the basic sot.
ellite vehicte for the Advanced Military Satellite
Programs, flight testing of the AGENA was accel.
erated when the DISCOVERER Program was
created, using the THOR/AGENA combination,
Because of its voilability, the Bell Aircraft
LRB1-Ba-3 rocket engine was selected for AGENA
Propulsion, and later modified to use unsymmet-
tical di-methy! hydrazine instead of JP-4 fuel.
Subsequent modifications resulted in the AGENA
“BY configuration, in which Propeliant tank
capacity was doubled and the engine modified
to provide single restart and extended burn capa-
bilities. The increased performance of this design
greatly enhanced the potential of the THOR/

for guidance and orbital attitude control was
developed to mest the critical orbital eccentricity
‘and attitude requirements for the programs in-
volved. Gas jets and reaction wheels are used
fo control attitude. Payloads may be installed
on the forward equipment rack or distributed
throughout the vehicle. The flight test program
also has been used to develop a recovery capa-
bility for a Payload capsule which is ejected
from the orbiting satellite,

The ABLE-STAR “pper stage vehicle contains an AJ10-104 propuision system which
is an advanced version of earlier Aerojet-Generai systems. In addition to Providing

increased performance capability, the system includes automatic starting, restarting, .

shutdown, ground control, coast period pitch and yaw control, and ground monitor.
ing systems, Propellants are fed to the thrust chamber by a high pressure helium
gas system, The thrust chamber is gimballed by electrical signals to provide pitch
and yaw control during powered flight. Roll control during powered flight is

Attitude control for coast Periods up to one-half hour Provided in the current design
can be extended by increasing the nitrogen supply.

AGENA

Prime contractor:
Lockheed Missite and Spoce Division

Engine manufacturer:
felt Aircraft Corp.

Length
“A" version 14 feet
“8” version 19.5 feet®

21 feet**
Diameter @ inches
Weight
“AS version 7,987 pounds
“8” version 14,800 pounds
“A" version ‘YURBI-80-5
“B" version XLRB1-B0-7*

‘XLRS1-Bo-9°*
Foot UDMH
Oxidizer (PNA
Geidence ‘eptical-inertial
Used ws second stege tor:
DISCOVERER (XVI & subs}
SAMOS (fight 4 and subs)

MIDAS fflight 3 ond subs)

- ABLE-STAR Vehicle

Contractor:

Asrojet-General

Height 14 feat 3 inches
Diameter 4 feet 7 inches
Weight 9772 pounds
Engine ASIO-104

with Restart Capability
Noszle Expansion Ratio—40,1

Fuel
Unsymmetricol Dimethyt Hydrazine

Oxidizer
Inhibited Red Fuming Nitric Acid

Guidance
‘STL Advanced Guidance System
Serroughs J-1 Computer

Used es second stage for:
TRANSIT 18, 2A, 28
COURIER 1A, 18


a
Pit

"firings, several of which were restart operations. An

additional primary objective of the program was the

. testing of the fuel pressure powered hydraulic system,

A8

also planned for initial use on the DISCOVERER XVII
flight.

© Testing of nozzle extensions for the XLR-81Ba-9
engine continued during the month at Bell Aircraft
and Arnoid Engineering Development Center (AEDC).
The extensions are designed to increase the nozzle
ratio from 20:1 to 45:1. Two successful firings of a
fitanium extension were conducted ot Bell Aircraft,
The extension has been sent to AEDC for additional
testing and evaluation. Stainiess steel test extensions
providing area ratios of 45:1 and 60:1 were test
fired with favorable results at AEDC during March,

Antennas

@ the manufacturing of two transistorized S-band
beacons was completed during the month. In order
to expedite delivery of these items, formal accept-
ance testing was started by LMSD representatives
on 14 March at the manufacturer's plant. One of the
beacons will be used for type testing and the other

will become the flight article for the first AGENA "B"
flight.

Recoverable Capsule

@ Intensive efforts were continued during the month
to obtain additional data on which reliability studies
and capsule recovery probability analyses could be
based. Sufficient component data has not been

* received from the capsule contractor (General Elec-

tric Co.) to establish firm reliability estimates. LMSD
reliability engineers are reviewing this problem with
General Electric. Co.

Ground Support Equipment

@ Countdown time will be decreased and fabrica-
tion and checkout procedures standardized by the

a

replacement of the present 200 pin umbilical system
with a 100 pin system beginning with the launch of
DISCOVERER XVII. The new system also will Provide
increased reliability by the simplification of check-
out procedures, and will reduce over-all manhour
requirements. Existing wiring in the vehicle is being
divided according to launch countdown or compo-
nent checkout function. The countdewn. wiring will’
make up the 100 pin umbilical system. The remaining
wiring will be led into test plugs for simplification of
component checkout. Blockhouse and trailer wiring
will remain unchanged. .

@ Revised testing and component selection proce-
dures for Fairchild programmers were completed dus-
ing the month and given to the contractor. The new
Procedures will expedite future deliveries of this
critically short component,

Facilities

© Interim measures to reduce personne! hazards
related to propellont storage and handling methods
at Vandenberg AFB launch sites have been incor-
porated until permanent facilities can be provided.
These measures include: improved lighting facilities
{including standby lighting power provisions), a water.
distribution and tanking system {including emergency
showers and eye wash facilities), a reinforced flush-
ing ond dumping pit for IRFNA areas, improved
water drainage around propeliont shelters, and

‘Improved methods of maintaining access roads.
“@ = Modifications to the Hawaiian tracking station,

in support of the NASA TIROS Project, were com-
pleted prior to the launch of that vehicle. Fly-by tests
were completed successfully and exercises run to
determine the time required to convert the rador
equipment from DISCOVERER to TIROS use. Change-
over time was determined to be ‘approximately 20
minutes.

“CONFIDENT:

WDLPM-4 200


Performance Graph — Boostes

0000 T T T 800 rw } |
000 — ATLAS/AGENA “A” ond “ge -| TT— ATIAS/As10-104/ABL 248 -|
AD ss/a -[
00 5
4
2” 10 N\
i ~
2000,
4, ‘\ 2
1000
800 200 5
1 °
600 &
“00 FY
34 us ae 40 rt
% 7
200 ATLAS/AJ10-101/ABL 248 73
! 2 3 4 5 600!
4
ATLAS/MERCURY a
Boy o 5
i. N
600
j= N\
3 200 . N
5 10 2 . N\

PAYLOAD WEIGHT x 1000 POUNDS BURNOUT VELOCITY—FPs x 1000

1, AGENA “A”— Polar Orbit 4. Lunar Probe
2. AGENA "B"— Polar Orbit 5. Venus Probe
3. AMR—90 degrees

WDLPM-4 200

!


WDLPM-4 200

MONTHLY PROGRESS—Project 609A
Program Administration

@ NASA has decided thai an early launching of a
modified, unguided SCOUT is desirable. This test
vehicle will be launched in late April and will consist
of the first stage, a dummy second stage, a live third
stage, and a nose cap of weight equal to the fourth
stage. The purpose of this launch is to obtain fight

experience with critical hardware items at the earliest ©

Possible date.

‘@ On 3 March, representatives of NASA, Chance-
Vought, Minneapolis-Honeywell, Aeronutronic and"

BMD met fo. establish a firm 609A schedule which
could be met regardiess of further NASA develop-
ment slippages. A definite schedule outlining dates
on which each agency must complete their part of
the program and its relationship to the launch date
‘of each fight was agreed on, This is the first time
in the 609A program that ¢ complete and detailed
schedule has been agreed to by all agencies con-
cerned.

© “The FY 61 Financial Plan for Project 609A, con-
sisting of data and documentation required to define
and justify the effort programmed for 61, was sub-
mitted to WDLPM on 21 March,

@ Arevised 609A Development Plan is being pre-
Pared by AFBMD for submission to Hq ARDC in Aprii.
This plan will cover both the development and Opera-
tional phases and will Present some program reorien-
tation based on commitments to support the Navy
TRANSIT Program.

© A transfer of obligation authority for $777,000-

was received from AFSWC for four 609A (2-3-5-6
configuration) vehicles. AFBMD was requested to
Procure four additional 609A unguided vehicles with
AFSWC funds and launch them with AFSWC pay-
loads during CY 1960. AFSWC has Hq ARDC

approval of this program, Recessary funding, and a
high priority requirement to make these tests this
year. AFBMD is taking steps to integrate the addi-
fiona! 2-3-5-6 vehicles into the Development Test
launch schedule.

Test Requirements Conference

@ This conference, attended by a total of 57
epresentatives from AFCRC, AFSWC, APGC, AFMDC,
‘Aeronutronic, BMC and. AFBMD, was held 1-2
March at AFBMD to collect, review, and discuss Pay-
loads planned for the 609A perational program,
The experimentat requirements for CY 1961, which
will be used to formulate the vehicle budget to be
Presented to. BMC, were of Particular importance.
Approximately 108 experiments (AFCRC 62, WADD
25 and AFSWC 21} were scheduled. AFMDC has
requirements for testing guidance systems, but it
appeared that the present 609A recovery system is

met to plan the approach for evaluating, validating,
assigning priorities, and deciding whether: or not
some experiments shall be referred to other Pro-
grams, such.as ATLAS Piggyback, Following this
review, Aeronutronic will combine the validated
compatible experiments into Payloads and the
required vehicle configurations ‘nd quanties will be
determined.

Technical Progress

@ The ABL X254A6 (third stage} engine has suc-
cessfully completed two test firings under simulated
altitude conditions at Amold Engineering Develop-
ment Center. These tests completed the prequalifi-
cation ‘test Program. The first flight engine will be
delivered in April. ,



ROLL JETS HEAT SHIELD
ENTRANCE HATCH an

ESCAPE ROCKET

HORIZON scanner —~//

~ ESCAPE ROCKET. PITCH @ YAW JETS: \
PARACHUTE
- —-/ @s
4a - AMS _.
JETTISON - f bE

> ROCKET OSIGRA
RETROGRADE:

~ WEIGHT AT SEPARATION APPROX 25 LBS

ORBIT INCLINATION 33 DEGREES
ORBITAL ALTITUDE (05~115 MILES (n)
3-18

HEAT SHIELD ABLATIVE
RECOVERY WATER OR LAND
Figure 1. Complete vebicle (top view) with

tatellise installed on ATLAS booster, Manned satellite
{bottom view) ibowine pilot Light postion, end detail sis of rire and posignade rockets eat oie
safety systems escape rockets,

WDLPM4 200 -CONFIDENTIAL- Li

ORBITAL CYCLES



PROJECT MERCURY

See
AFBMD Responsjbitity
in support of
PROJECT MERCURY
NASA HS-36
includes:
HH

Design, engineering studies Launch support

Equipment modification Trajectory data

Hardware fabrication . Missile allocation
Flight scheduling

Provide fourteen (14) Modify boosters for NASA Prelim. Launch, control and define ralectorios
ATLAS boosters, inary research and manned orbital of booster-capsule Vehicle up to, and
. flight and safety objectives, induding, injection into orbit,

Table 2. AFBMD responsibilities ie suppors of PROJECT MERCURY,

MONTHLY PROGRESS—Project MERCURY on schedule, This launch is scheduled for late May
Program Admin he or early June from Atlantic Missile Range complex

“@ A revised HS-36 Development Plan for a 14 mis- Qpproximately 27 April.
sile/13 taunch, 51.5 million dollar ATLAS/MERCURY
Program has been approved by AFBMD and for- Technical Progress

warded to NASA, The funding breakdown is as g As @ result of NASA revising their test plans and

follows: . elimating the C-band radar beacon from on-board
Dotaas Capsule equipment the 50D ATIAS booster will carry
TSCAL YEAR Un Militens) @ General Electric impact Predictor.. The C-band
1959 8.761 beacon would have been sufficient to satisty range
1960 14.910 safety requlremants as a back-up to the Azusa impect .
1961 19.979 Predictor carried aboard the booster.
1962 7.854 © Aconcerted effort has been made to save weight
. TOTAL 351.504 in both booster and capsule design. One result of
Flight Test P, this effort isa lightweight telemetry system, identified

© Test preparation ‘for the second launch in the tion in the booster. It fs anticipated that this change
ATLAS/MERCURY flight test program is Proceeding will be incorporated on the ATLAS booster,


OT cost

fives. The designation and purpose c of each configu-
ration is as foltows:

Readout:
E-1—Component Test Payloads
E-2—Steerable Reconnaissance Payloads
{with 20-foot ground resolution)
Recovery:
E-5—High Resolution, Steerable, Recoverable
Payload [with 5-foot ground resolution)

Payloads

@ B11 Payloads—Subsystem testing of the first
E-1 payload has been completed successfully and
payload compatability with the AGENA vehicle has
been established. Systems testing is now underway.
Both the E-1 and F-1 payloads will be included in
stotic hot firing tests at Santa Cruz Test Base. Accept-
ance tests of the second E-1 payload were started on
17 March at Eastman Kodak, with delivery to LMSD
scheduled for early April. A 48-hour operating test
performed during the modification and checkout
procedure revealed a high voltage power supply
deficiency. A back up development has been initi-
ated on this item.

@ £-2 Payloads—Fabrication and assembly of
components for the first two E-2 payloads is nearing
completion. Component testing has been underway
since mid-March and is scheduled for completion by
mid-April.

© E-5 Payloads—Design review meetings are
being held on the E-5 system functional requirements
to establish firm definitions of camera design, cap-
sule configuration and recovery sequence. Prelim-
inary design of the 66-inch lens camera continues
at the ftek Corporation. A contract was let to AVCO.
Corporation for the recovery capsule ablative shield
plus other required heat shielding, and ‘preliminary
work has been started. Several recovery capsule con-
figurations are being evaluated and wind tunnel and
ballistic range tests are being performed. Both over-
land and over-water recovery capability will be
included.

Ground Support Equipment

@ All payload support equipment functioned suc-
cessfully during a subsystem test with the E-1 payload
mounted on the 40-inch optical collimator. Satisfac-
tory operation was also experienced during a sub-
sequent preliminary system test with the payload
mated to the AGENA vehicle.

MMT ELEN Hinesse

Ferret Reconnaissance System

© Ferret Reconnaissance System payloads are being
developed in a minimym number of configurations.
The designation and purpose of each configuration
is as follows:

F-1—R&D Test Payloads

F-2—Digital General Coverage Payloads

F.3—Specific Mission Payloads—Analog Presen-
tation

F-4—Technical Analysis (study stage only)

Payloads

@ = F-1 Payloads— Subsystem tests of the first two
F-1 payloads have been completed. The first F-1 pay-
load has been aligned with the first E-1 payload and
the dual package installed in the AGENA vehicle.
Systems testing of the complete installation were
started late in March.

@ = F-2 and F-3 Payloads—Design and modifica-
tion of the reoriented F-2 and F-3 systems ore in
progress. Delivery of the first F-2 flight article pay-
load is scheduled for October and the first F-3 is
scheduled for 1962 delivery.

@ F-4 Paylouds—Studies of the F-4 System are
being continued at Airborne Instruments Laboratory,
based on 1965 intelligence requirements.

Ground Support Equipment

© Mechanical inspection by the Air Force of the F-1
data conversion equipment was completed on 22
March. This equipment will be installed in the Satel-
lite Test Center, Sunnyvale, California. The checkout
equipment for subsystem testing of F-1 payloads was
shipped to LMSD on 25 March. This equipment will
be instalied in the missile assembly building at Van-
denberg AFB,

Communications and Contro! Equipment

© The UHF ground system is being dismantled by
the contractor for shipment to Vandenberg AFB.
Included in the system are the angle tracker, com-
mand tracker, command transmitter, and the tracking
and acquisition system, The 60-foot tracking and
acquisition antenna will be. dismantled and trans-
ported by police escorted truck to the dock at Moffett
Field, California. It will then be loaded on a barge
for shipment to Vandenberg AFB.





CONFIDENTIAL

” FIBERGLAS
ATTACH RING

UPPER WINDOW

REACTION

th iM MN

NOZZLES

ASSEMBLY

WINDOW t S Da =

SHIELD

WO

FLOTATION
BAGS ACCESS

‘AF TERBODY

Figure 3. McDonnell shell to be carried on second
ATLAS/MERCURY flight s01s.

WDLPM4 200 ~CONFIDENTIAL — Ls


ABLE-4 THOR (PIONEER V)

@ On 11 March 1960, at 0800 EST, the ABLE-4
THOR interplanetary probe flight test vehicle was
launched from Atfantic Missile Range (AMR) Pad 17A,
‘Twenty-five minutes later, following excellent per-
formance by all three stages, satellite vehicle
separation occurred and the PIONEER V satellite was
placed in an elliptical orbit about the sun, At this
time, all aspects of this flight continue to produce
the desired results. Performance of the three booster
stages, communications level, and information from
Payload experiments have far exceeded any previ-
ous achievements in space research. ABLE-4 THOR
was the third fiight in the ABLE Space Probes Pro-
gram. The vehicle and payload were similar to those

tested in the highty successful ABLE-3 earth satellite
(EXPLORER VI).

@ = The ABLE flights began in March 1958 to provide
fight test support for development of ballistic missile
re-entry vehicles over intercontinental trajectories.
Successful R&D flights led to the use of the THOR/
ABLE vehicle, with a third stage added, for deep
space and interplanetary probe missions. This has
Proven to be an exceptionally capable and reliable
vehicle. No other comparable ballistic missile hos
ever attempted the payload weight handled by this
vehicle.

@ The high degree of success achieved resulted
from close and effective administrative and technical
coordination among the National Aeronautics and
Space Administration (NASA) and the Air Force
Ballistic Missile Division/Space Technology Labora-
tories team.

Trajectory—ABLE-4 THOR (PIONEER V)

E7





SW3LSAS 3D dS

systems

SATELLITE

DISCOVERER
SAMOS
COMMUNICATIONS

MIDAS



CONFIDENTIAL

@ = Receiver—The poyload command receiver is a
transistorized double-conversion, phase-lock-loop
receiver which produces a coherent output at 2/17
of the receiver frequency. It can be operated with
an: effective noise bandwidth of either a 25-cps ora
40-cps bandwidth. The receiver operates continu-
ously and, since its bandwidth is less than the
frequency uncertainty of the received signal, it
repeatedly sweeps over a range of 30 ke searching
for a carrier. The sweep period is 10 seconds for the
wide band and three minutes for the narrow band.

When the receiver acquires and locks on a signal
from the earth, the sweeping stops and the receiver
can then accept any of eight possible commands, as
follows:

1. Transmitters “off.”

2. Five-watt transmitter “on‘ at 64 pulses per

second.
3. Five-watt transmitter Yon“ at 8 pps.
4. Five-watt transmitter “on"' at 1 Pps.

5. Stage three separation.
6. Receiver narrow band,

7. 150-watt transmitter plates on.
8. 150-watt transmitter filament on.

© Signats from the earth to the satellite are
transmitted by using a high-power corrier, phase-
modulated with a 512 cps subcarrier. Amplitude
modulation of the subcarrier with a coded train of
13 pulses provides the required information to the
satellite command decoder. The receiver weighs four
pounds, occupies 1300 cubic inches, and draws 1.5
watts at 16 volts,

@ Range and Range Rate Information—Satellite

range is measured by frequency modulating the _

ground station transmitter by low-frequency
sinusoids. A modulating frequency of 16 cps, when
adequately smoothed, provides range accuracies of
better than 100 miles. Additional modulation fre-
quencies are used to resolve other signal ambiquities.
Angle tracking is performed by nodding the ground
antenna alternately in elevation and azimuth as it
aims et the payload. Angular accuracies of about 0.2
degrees are possible with the four-station network.
Trajectory smoothing by the ground network can
Provide velocity and angular measurements within
0.01 degree and 0.1 ft./sec after the first few days
of tracking.

@ = Range.rate is being measured to accuracies of
1 ft./sec by extracting the two-way doppler fre-
quency shift between the transmitted and received

CONFIDENTIAL

signal, after correcting for the frequency offset
introduced by the satellite transponder.

Telebit System

© This digital telemetry unit, developed by Space
Technology Laboratories accepts both analog and
digital inputs from the various payload experiments
and satellite performance transducers. The con-
verted information appears. at its output as a binary
coded subcarrier (1024 cps) which then phase
modulates the signal of the transmitter. All satellite
data is contained in seven separate “words which
are sent to the ground. Six of these are information
on experiments. The seventh is subcommutator meas-
urement information consisting of: paddle tempera-
tures forward (1) and oft (2), transmitter heat-sink
temperature {3}, battery temperature (4), internal
ambient temperature (5}, converter heat-sink tem-
Perature (6}, battery voltage {7}, and receiver phase
error (8). - ‘

@ The binary output of the system occurs at a
synchronous rate and is composed of repeating sets
of frames of words. Eight words per frame are used
for the PIONEER ¥ payload. One word of each frame
is used for frame synchronization and is read out as
all zeros. The other seven words are coded with the
digital representation of the input information. Each
word contains 12 pulses. The first two pulses of
information words represent xero and one to pro-
vide a word synchronization. The other ten pulses
contain the experimental data or vehicle.measure-
ments in binary number form.

@ = The conversion of analog to digital information
is accomplished with a digital ramp ard a voltage
comparison ‘circuit. The counting Is actually - per-
formed in a shifting accumulator, just es for digital
data,

Power Supply

@ Solar Cells—A total of 4,800 boron-diffused
silicon solar cells are carried on the satellite's four
tolor paddles. Because of attitude and spin consider-
ations, however, only about 1,000 of the cells are
receiving solar energy at a given time.

@ Batteries—the satellite storage battery consists
of two packs of 14 nickel-cadmium cells each. Part
of the power developed in the solar cells is consumed
immediately by the experiments and the receiver,
and the charging rate of the battery is lessened by
this amount, The 5-walt transmitter, drawing a large
Portion of its power from the batteries, can be oper-

‘WDLPM-4 200


WDLPM-4 200

ASPECT
INDICATOR

BATTERIES

COMMAND BOX
DIGITAL
OEcon!

SEARCH COIL
MAGNI

LOGIC
(UNDERNEATH)

TELEVISION
LOGIC CIRCUIT

COMMAND RECEIVER

SOLAR CELL
CURRENT MONITOR

FLUX GATE
MAGNE TOME TER.

THOR-ABLE II PAYLOAD

SUBCARRIER OSCILLATOR
A
MULTIPLEXER "B"

INJECTION ROCKET

HEAT SINK

SW TRANSMITTER

PROPORTIONAL,
COUNTER
+ {CHICAGO)

STATIC
CONVERTER
toe Mc
TRANSMITTER
BATTERIES

SUBCARRIER OSCILLATOR
AND
MULTIPLEXER “A”

OIGITAL
TELEMETRY

(TOP VIEW)

Figure 4, ABLE-3 payload (sop view),

and orbit were essentially as Predicted with devia-
tions in apogee and perigee occurring on the mare
than nominal side. The payload was the most sophis-
ticated to have been piaced in orbit by this nation
at the time and contained Provisions for conducting
13 experiments in space environment and propaga-
fon. A wealth of valuable data was obtuined from
satellite telemetry until the ‘ast transmission was
received on 6 October. It is believed that the satel-
lite, while yet in orbit, is incapable of generating

sufficient power for transmitting 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 IRBM. 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 . ze


Project MERCURY represents the transitional threshold . viding 14 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 vehicies through injection into orbit.

manned satellite into orbit about the earth, and to The division of responsibilities for this Program is given

effect a controlied re-entry and successful recovery in Table 1. Specific details of AFBMD support are given

of the man and capsule (Figure 1). Unmanned ICBM in Toble 2,

trajectory and near-orbitat flights, and unmanned :

orbiting flights will be used to verify the effectiveness Major contractors Participating in the AFBMD portion -

ond reliability of an extensive research program prior of this. program include: Space Technology Labora-

to manned orbital flights (Figure 2). The Program will tories, systems engineering and technical direction;

be conducted over a Period of nearly two years. The Convair-Astronautics, modified ATLAS boosters; _

initial R&D flight test was ‘accomplished successfully GE/Burroughs, ATLAS guidance equipment; and Rock-

in September 1959. The total Program accomplish. etdyne, engines. All of these companies also provide

ment is under the direction of NASA. The primary Special studies and engineering efforts peculiar to

responsibility of AFBMD to date consists of: (2) pro- meeting Project MERCURY requirements. 7
a a

LEGEND ; — —€ -

RESPONSIBILITY:

AFBMD =e ‘

NASA amen

ESCAPE ! ‘
ROCKET

' / @)BALLIsTic TRAJECTORY
1 / TO SIMULATE REENTRY CONDITIONS

7 .

7 NEAR ORBITAL INSERTION, IMMEDIATE
/ REORIENTATION, DECELERATION, REENTRY
/ ATLANTIC RECOVERY
4

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

ee comme

Figure 2. Flight test srajectories for Project MERCURY, defining specific objectives, ‘Thejectory C
‘ tad ie Path ofthe fal (manned) fight. The point ot which APBMD ene tater ibilisy

a CONFIDENTIAL WDLPM4 200



MONTHLY PROGRESS—MIDAS Program
Flight Test Progress
First MIDAS Flight Test

© As reported last month, the first MIDAS flight
test vehicle was launched 00 26 February from the
Atlantic Missile Range. Analysis of flight data indi-
cate that an event occurred at the approximate time
of first and second stage separation which’ damaged
the AGENA vehicle and prevented the attainment of
@ successful orbit. Significant events occurred in the
following sequence:

@. Final countdown was smooth and liftoff normal.

'b. ATLAS performance was normal up to the time
of ATLAS/AGENA separation,

¢. AGENA seporation command was received at

rat

T plus 258 seconds, following ATLAS vernier
engine cutoff.

d. AGENA telemetry signals were lost temporarily
at T plus 258 seconds, then received inter-
mittently and with highly degraded quality
until approximate T plus 325 seconds,

@. ATLAS telemetry, at reduced signal strength,
continued until T plus 989 seconds.

f. Radar tracking data confirmed that the AGENA
did not attain orbit.

Second MIDAS Flight Test

@ The second MIDAS flight test has been resched-
uled to 20 April. The vehicle will be launched from
Atientic Missile Range pad 14. The MIDAS infrared
payload will be placed into a 261 nautical mile
orbit. Orbital time will be 94 minutes (nominal) and



PIONEER V
PAYLOAD

(top view)

lonization Chamber
and Geiger-Muller Tube

: Heat-Sink—
150-watt Transmitter


CONCEPT

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

lated ground

three strategically located readout stations,
The readout stations transmit the data
directly to the MIDAS Control Center where
it is processed, displayed, and evaluated
(Figure 5.) Hf 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 retaliatory and defense
agencies.

determine when an actual atiack
been launched. The decision 1

other using ies having a corre-
agencies 4



Time Time
(sec) Stage Event sec) Stoge Event
X+01 1 Liftoff switch activates X + 2300 Il Jettison nose fairing
Programmer storts Times varyin fl Stop pitch program
Gyros uncaged accord. with Pitch command
Nt Umbilicals eject trajectory. Yaw command
Arm destruct initiator X — 429.0 It Stop pitch program
X+2 { Roll program initiated X— 4415 Ml Engine cut-off signal
X49 1 Roll program complete ‘Switch pneumatic coast control
x+10 __ Pitch program iniiated tem on pitch, yow and rol
Ist step pitch rate As required X- 4710 a Tern off hydraulic power
X+25 \ “Ghd step pitch rate for trajectory X ~ 480.0 It Initiate coast phase pitch program
x+70. | “rd step pitch rate ond detailed in XT10868 Stop coast phase pitch program
x +90 | “Ah step pitch rate DTO X= M86 It Stor hydraulic power
‘Autopilot gain change X— 14886 il Engine restart fire signal
Programmer armed Uncage accelerometer
x4 190 | Pitch program complete X— 1491.6 Cage coast pneumatic control system
X+ 152 |” Main engine cut-off (MECO] X— 1500 _Arm TPS cut-off probe (back-up)
Circuitry armed Arm oxidizer probe (back-up)
x4 1035 1 MECO bockup ormed : Arm spin and seperation mechanism
x4 1670 1 MECO x 7 1504.0 il Engine cut-off signal
X+1670 Wl Stort programmer \Urcage coast pneumatic control system
X+1700 i Engine Fre signal Short spin table
Uncage thrust chamber Start timer on spin table
in pitch and yow X — 1506.0 WW Engine cut-off (back-up)
Uncage high thrust roll jets Start spin table (back-up)
Uncage gyros X — 1507.0 Remove spin table bolt power
X + 170.85 Blow seporation bolts X — 1526.0 {| Blow separation bolts
X+ 1760 Nt Start pitch program Activate separation actuators
Payload separation occurs
Sequence of Events—TRANSIT 1B, 2A end 2B
P4 iP WDLPM-4 200



THOR/ASIO-42/Aa1 248 3000 THOR/ANO-101 7481 248
2000
q 1000
t 800
600
L 1 Fr 1
400
L > L )
200
A A 1 A f 1 ' F ' A A
2 4 6 8 10 12 2 4 6 8 10 12
THONANO-104 3000 | THOR/AGENA “A ola =a
2000
i 1000
L r 4
800
a r 5
600
L 1 L
409
L 2
3.
200
F F F F : 1 \ A 1 \
2 4 é 8 10 12 2 4 é 8 10 12
Paylood Weight x 100 pounds Paylood Weight x 100 pounds
1, Polar—AMR or VAFB 4. VAFB—AGENA “3” (XLR81-Ba.7)
2. AMR—90 degrees 5. VAFB—AGENA “p" (CXLR81-Ba-9)
3. VAFB—AGENA “at
wn a ee



The initial visual recownaissence program will use con- stations thru a data link. Ground electronics

techniques with astomatic film proc- vert the si

santo photo image form,

Figare 3.
Schematic of SAMOS system
in operational orbit, When the
cele oer wnfrendly sor
ritory the sensing equipment is
turned on (Information gath-
Find een cd Ieaves an
friendly territory the sensin,
ipment is turned off od
the sensing date is processed
(Information storage). When
she vebicle comes within range

of s receiving station,
dy will be read-ont spon
pened oreeeiting and

ital to asing agencies.
This process is continously
Cpeeaied daring mseful life-

time of the

swith « capability
‘TV-type electronic image readout to growed of rcoley abies far eee ‘


PROJECT MERCURY
NASA Responsibility AFBMD Responsibility
Booster-Capsule integration |
Capsule Booster
Ground support ATLAS “D"
Recovery -
Modified
|
Flight 1 Flights INew guidance antennae
| . 2-12 Thin skinned tanks
NASA it ‘Shifted tank bulkheads
Laboratory Flight Light weight thrust structure
Capsule Capsules Capsule separation mechanism
(McDonnell) interface electrical connections
LOX dome protection
Re-entry Programmer Modified
heating, Range safety destruct time delay
dynamics, |Abort-sensing system
recovery

Unmanned non-orbital fights

Manned and unmanned orbital fights

Launch from AMR

Earth track 33°N/33°S/Lat. (for orbital
flights}

3-18 orbits

Orbit capability 28 hours

Orbit weight 2500 tbs,

Orbit altitude 105-115 nm.

Automatic pilot safety equipment with

manual backup

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

Integrated countdown—inciuding pilot insertion
and capsule countdown. AMR launch Stand 14—
adapted for pilot and capsule.

GE ground guidance equipment modified to
Provide: {a real time data, {b) additional discrete
command signals during powered flight.

{a} Ballistic trajectory—108°
(b} Orbital trajectory —072°

Table 1. Outline of NASA and AFBMD responsibilities ia PROJECT MERCURY.

“CONFIDENTIAL-


CONFIDENTIAL

will be approximately 24,869 miles per hour.
PIONEER Y will have an orbital “year” (time period
to complete an orbit) of approximately 312 days.

Period 311.64 days

Time to perihelion 151.96 days

Eccentricity 0.104 .

Semimojor axis 0.889 astronomical
unit

inclination ecliptic 3.35°

Longitude to first —10.29°

ascending node

Arguement of perigee 2.58
Table 2. ACTUAL ORBITAL PARAMETERS

@ Performance figures for PIONEER V, as af 22
April, are as follows:

© Days since founch .............000., 42
@ Miles from earth . . + 4,915,000 n.m.
® Velocity {in relation to earth) 5,126 o.m./hr.
@ Telemetry reception rate, (see
communications section) ...... +++ 8 pps
© Temperature {solar ceil
paddles) .........50 degrees Fahrenheit
Ground Station Tracking
and Data Acquisition

@ = Ailantic Missile Range—Al\ launch and pow.
ered flight tracking operations performed by this
station were completed satisfactorily. The station was
able to track the satellite only for the first five days.

© = Manchester, England — The 250-foot antenna
at this station acquired the ABLE-4 THOR velilcle 15
minutes after launch. Reception has been. excellent
throughout the flight to date, in spite of several
equipment failures; all of which have been repaired.
Only a small amount of transmitter noise has been
encountered. The parametric amplifier is performing
quite satisfactorily and has been used throughout
the flight. It has the usual noise figure of 2 db.

@ South Point, Hawail — Telemetry reception by
this station has been second only to Manchester.
Because of a transmitter failure, which has been
repaired, tracking was not accomplished for two
days. Late in April the reception rate had dropped
to 1 pulse per second {from the 5-watt transmitter}
‘and quality had decreased to “garbled.” A decrease
in transmitter noise was realized by the installation
of the 3% Jnch styrofiex cable and the welding of
the screen of the antenna reflector. Parametric ampli-
fier performance has been satistactory.

CONFIDENTIAL

@ Singapore —This station was able to transmit
commands to the satellite for the first two days in
orbit and to receive telemeiry for.the first five days.
All equipment operated, satisfactorily.

© Millstone Hill, New Hampshire — Reception at
this station was very excellent via the regular 400 mc
radar feed attached to the 85-foot antenna, Because
of Its lobing capability, Millstone was able to give
very accurate satellite angle information.

Satellite Vehicle

©@ The ABLE-4 THOR satellite is an approximate
spheroid, measuring 26 inches in diameter at its
equator. Four solar paddles are evenly spaced about
its equator. The paddles were folded downward
during first and second stage Powered flight. At the
time of second stage shutdown the paddles were
released and erected in their Permanent positions,
prior to spin up and firing of the third stage engine.
Each of the solar paddies Is offset in angle about 30
degrees relative to the vehicle spin axis to permit
@ relatively constant number of solar cells to be
‘exposed to the sun despite the changing attitude
of the satellite. .

@ The satellite weight is 95 pounds, distributed as
follows:

Structure .. + TN

Platform and wiring .. - 74

instrumentation ...

Experiments .

Batteries ...... tees .

Paddles and solar cells .......... 10.2

Temperature control material ...... 0.8

Dynamic balance weights... 17
Satellite Instrumentation

@ Communications — Communications equipment
aboard the PIONEER V is by far the most powerful
‘ever carried on a missile or space flight. information
will be sent te earth from both the 5-watt and-150-
Watt transmitters. A high sensitivity receiver permits
satellite reception of earth transmitted commands.
When interconnected coherently, both range and
range rate information may be obtained to permit
accurate tracking of the satellite.

@ Transmitters — Satellite transmissions to date
have been performed by the 5-watt transmitter. It is
intended that this procedure will be followed until

WDLPM-4 200


K2

MONTHLY PROGRESS—ABLE-STAR Program
e All work related to the development of the
ABLE-STAR (AJ10-104) propulsion system os directed
by ARPA Order No. 95-60 ‘was completed during the

month. A final report to ARPA will be submitted by
AFBMD in April.

© Initial fight test of the ABLE-STAR vehicle is
scheduled for 13 April, as the second stage for the
TRANSIT 1B. Additional ABLE-STAR flight tests are
Programmed at present for TRANSIT 2A and 28, and
COURIER 1A and 1B,

rine are

Woy

WDLPM-4 200


WDLPM-4 200

HzO. JETS

— 30°

(IMPROVED)
XM33E5
(ALT. NOZZLE)
Hy Oo, JETS

j—— 31"

33KS 120,000
JET VANES AND
AERO FINS

—————$
PROJECT 609A

Hyper-Environment Test System

PROGRAM DESCRIPTION—The Hyper-Environment
‘Test Program (609A) ls divided into R&D and Opera-
flonal Phases. The R&D phase will be weed to develop
‘and fight test vehicles capable of carrying 50 to
1,000 pound payloads to altitudes of 200 to 7,000
miles. The Operational phase will use this standard.
ized vehicle 10 permit the sconomical performance
of flight test experiments in support of scientific
research and advanced military space system
programs,

iconomy—Reliabllity—Versattilty—in this order
Of emphasis ore the three significant guides to pro-
gram accomplishment. ECONOMY is being achieved

ABL X248A6—_
ABL X254 A2

THIOKOL

SERGEANT Nigure 1. Tbvee sariaions
. HM3BES : Freee! 6094 webicle

L. NOZZLI Smcemint the sisi
JET VANES AND - cereale cf the trogen.

AERO FINS



A. SECOND STAGE—ANLE-STAR {As10-104)
1 Thrust of altitude 9030 pounds
Specific impulse {vac} 278 seconds
Total impulse (min) 2.3 X 10° tbs/s0c
Burning time 294 seconds
Propeliant Uquid
H B. rest stacts—mon msm
Thrust (3.1) 151,500 pounds
‘Specific impulse (s. 1) 248 seconds
4 ‘Specific impulse (roc) 287 seconds,
. Burning time 158 seconds
Propetiant Liquid

UPPER PAYLOAD *

PAYLOAD
h (36-INCH DIAMETER)

2 PAYLOAD
<A SEPERATION PLANE
PAYLOAD
. SPIN MECHANISM
~Equiement

COMPARTMENT

‘
ets
if l . | pRoPULSION SYSTEM
i

54.79 DIAMETER

1. Provide accurate navigational reference informa-
tion for POLARIS launches, :
2. Precise determination of satellite Position by
measuring the doppler shift of satelite transmitted
radio signals,

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

4. Acquire additional geodetic and geographical
data by precision tracking of the orbiting sateliite.
Flight Vehicles TRANSIT 1A consisted of three
stages as shown in Figure « TRANSIT 18, 2A
and 2B are two-sige vehicles as shown in Figure.
Launch Plans All vehicles will be launched from

Atlantic Missile Range pad 17A or 178. Launch
azimuth for TRANSITS 1A and 1B is 44.5 degrees
and for TRANSITS 2A ond 2B, 140 degrees.

1 TRANSIT 18, 2A and 28

GIMBAL ¢

TS stase 1/11

SEPARATION PL ANE

‘STAGE I

% 20 INCH DIAMETER
TRANSIT 2A & 2B ONLY

Powered Flight T ‘The powered flight
trajectory for TRANSITS 18, 2A and 28 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
opproximately 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 agate
which permits the insertion of data only when the
gate has been opened at a previously scheduled
fime. 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 end 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.

STAGE I/I. ADAPTER

GUIDANCE COMPARTMENT

ra GR CODEN T throes a


D2

more advanced phases was teplaced by a single .

integrated effort to which the code name ADVENT
‘was applied, ADVENT calls for an R&D program for
@ 24-hour globo! communications satellite system.
The feasibility of placing a satellite in a predeter.
mined position in a 19,300 mile equatorial orbit must
be demonstrated. The satellite must be capable of
Providing worldwide communications on a real time
basis at microwave frequencies with a high channel
wide bandwidth capacity. Amendment 5 also
requires the design of a single final stage vehicle for
microwave equipment compatible with launching by

either AGENA “B" or CENTAUR second stage
boosters.

The ADVENT program, as defined in amendment
5, will consist of the following flight tests:

. Four ATLAS/AGENA “BY flights, nominal 5,600
Nautical mile orbits,

b. Seven ATLAS/CENTAUR flights, launched from
the Atfantic Missile Range into 19,000 mile
equatorial orbits. :

c Three flight tests, using payload space on the

last three NASA ATLAS/CENTAUR R&D flights.

Figure 2. Artist's concept of proposed microwave sommasications,



WDLPM-4 200

Figare 1.
ing of AGENA "BY
bropalsion system.

The AGENA vehicle was originally designed by the
Air Force for use as the basic satellite vehicle for
the Advanced Military Reconnaissance Satellite Pro-
gram (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 inertial 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
fo 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-methyl hydrazine {UDMH) fuel instead of
5P-4, Accelerated flight testing of the AGENA ve-
hicle and its subsystems became possible when the
DISCOVERER program was created, based on the

low cost and early availability of the THOR IRBM,
ond @ 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 initiated to develop the AGENA "B”
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 XLRB1-Ba-7, A.

subsequent version, XLR81-Ba-9 (Bell Aircraft Model
8096), with a nozzle expansion ratio of approxi-
mately 45:1 will further increase Performance capa-

Il



SLD3fOad IDwdS

projects

COURIER
TOS
AGENA
ABLE-STAR
MERCURY
609A

ABLE
TRANSIT



Solar Paddles...

In folded {launch} position (left) . . . in extended
(orbital) position (below).

Servicing...

+ + « Flight vehicle with fuel (left}
and oxidizer (right above).

zi CONFIDENTIAL wou


Figare + Mees ronceps fh, pas

o_o me oe” Oo ome

“an —"
Weight—Inert 1,370 1,600
Impole Propelionts 6,560 13,100

-_ om
3
i

1
i
i

Pyrotechnics CA

GROSS WEIGHT (be) 707 14200
Engine YLRSIBe-5 XiR81-B0-7

Threst, les. tree) 15,000 15,000 Ss
Spec. Imp., sec. (vac) a7 a7
120 240
Ne Yes

u
3
fi

SOLAR
| ARRAY
BEACON OESTRUCT
ANTENNA:

4
H

SHIELD ANTENNA = SEPARATION ANTENNA COMPT. TANK TANK

SOOSTER—ATLAS ICOM
NOTE: AGENA “A” configurotion except

for solar paddies (AGENA:“B” only}. no eo
: Onidizer (LON ‘17a300
GROSS WEIGHT (ibe.) 262,300 ;
. Engine—MA-2

Sustoiner 82,100
Spec. Imp. (see. vac} Roce 286
Sestoiner 310



‘WDLPM-4 200

Ready for Launch...

++. from Atlantic Missile Range
launch pad 174

E17



A, THIRD STAGE—X-248 (Allegany Ballistic Lak.) os
Thrust at altitude 3150 pounds A

Specific impulea (voc) 250 seconde +—

‘Total impulee 146,400 Ibs /s0¢

Burning Time 37.5 seconds

Propellant Solid 8 é
§. SECOND stAGE—As10-42 (AcrojerGeneral) _ | i

Thruat at altitude 7700 pounds "

‘Specific Impuise (voc) 271 seconds ;

Totol impulse (nin) 870,000 tha/s0e é

Berning time 115 seconds a

Propaitont Uguid

C, FIRST STAGE—THOR RBM

Thrust (6.1) 151,500 pounds
Specific impelse (s.1,) 248 seconds
Specific impuise (vac) 287 seconds
Burning time 158 seconds
Propellont Uquid

AY)

Figure 1. TRANSIT 1A shree stage flight vebicle,

Rake: + ~

TRANSIT IA lewached from Aslaxtic 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
add TRANSIT 1B, 2A and 2B-flights. The program is
currently authorized by ARPA Order No. 97, which
ausigns AFBMD responsibility for providing the
booster vehicles, integrating Payloads to the vehi-
cles, and flight operations from launch through
attainment of orbit, including communications to the
tracking and data handling facilities. Payload ond
tracking responsibility has been assigned to the
Figure 2. Cat-aoay dressing of TRANSIT IA payload (NAV 1). aha ot Ocinonee. Apeted Physics Labora-

‘WDLPM-4 200 F-l





CONFIDENTIAL

Preparing the Stages...

Second Stage Checkout . .

Systems checkout of second stage being performed
in hangar at Attantic Missile Range, Photo

sl ABLE vehi¢le mounted on transporter;
mobile checkout vans in background.

Third Stage—
Payload...

Third stage, with test payload
installed, during stress testing of
installation. Extensive iast minute
vibration tests were conducted on the
third stage with highly successful
results being obtained,

CONFIDENTIAL wou 0




A

Figure 4. Preparing the payload
intllaiog guid ihe foleed ee

wad

a)

Fe



i° {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.

*  (f) Conducting bio-medical xperiments with mice and

; small primates, incl ing injection into orbit, re-
: entry and recovery,

.

Early tests confirmed vehicle flight and satellite orbit
capabilities, developed system reliability and predict-

data acquisition requirements. Subsequent flights are
planned to acquire scientific data for design of
military reconnaissance Payload compon-

ents. Typical data Gathering objectives indude: cosmic
‘and atomic radiation, magnetic field, total electron

| density, auroral radiation, micromsteorite measure-
ment, Lyman alpha from space (or stars), solar radia-

tion, and atmosphere density (drag) and composition,

A4

- IMPACT
AREA

SEPARATION
COMMAND

PALO ALTQ
a

oN.

Alaskan and Hawaiian fy

seberation point, and impact area, Inset at left shows
vebicle attitude required to effect capsule re-entry
following ejection, i

A world-wide network of control,

Figure 4, Recovery sequence, showing tracking by
‘ ‘ bick

Control Center.

Telemetry ships are positioned @s required by the
specific mission of each flight. Figures 2 and 3 show
@ typical launch trajectory from Vandenberg Air
Force Bate, and figure 3 shows schematically 2 typ
ical orbit. An additional objective of this program is
the development of © controlled Fe-entry and

— (CONTROL

VOICE AND
DATA CABLE

16 th ORBIT



CONFIDENTIAL.

Launch... Launch

©@ Countdown was initiated at 0300 EST, on 11
March. Final missile preparations, including ordnance
-Provisions, were compieted by 0430. The final _
checkout tasks (advanced guidance system and pay-
load) were completed at ‘approximately 0540. The
one-hour hold programmed into the countdown at T
minus 35, was started at 0625. Terminal countdown
wes resumed at 0725:30 and a time synchronization
made. The blockhouse was secured at 0736. Launch
Grea wos declared to be clear at T minus 27 seconds.
Blockhouse vents were closed at T minus 6 seconds.
Countdown remained on schedule untit launch.

Powered Flight

@ Liftoff occurred at 0800:07:1 on a launch azi-
muth of 90 degrees true. Gross weight at liftoff was
113,675.01 pounds. Performance of ail three stages
was excellent and separation occurred as planned.
Burning time for the THOR and third stage engines
was within one second of nominal. The Aerojet
AS10-101A second-stage engine cut off within one
tenth of one second of nominal burn time. The third
stage separated from the terminal vehicle at 0825:15
EST. Velocity, at the time of third stage separation
was approximately 36,480 feet per second, over 800



“CONFIDENTIAL

Permit an accurate determination of the length of
the astronomical unit. This unit, the semi-major axis
of the earth's orbit, is a basic astronomical constant.
It Is the unit for determining the absolute distances
of solar system bodies from each other; it is the basis
for the parsec unit of astronomical distances; its
magnitude, and the derived value of the mass of the

sun, ore fundamental to all space trajectory calcula-
tions. Depending upen the length of time that
PIONEER V will be tracked, a determination of the
Astronomical Unit within an accuracy of 0.03 per-
cent or better is anticipated, thus providing a cross-
check of the AU valve as determined optically and
by reflecting radar signals off the plonet Venus.

WDLPM- 200



Monthly Progress-—DiSCOVERER Program
Flight Test Progress
@ DISCOVERER Xi was launched from pad 5, Van-
denberg AFB, at 1230 PST on 15 April. The count-
* down proceeded very smoothly despite high winds.
No technical holds were encountered, but the launch,
scheduled for 1100 hours, was delayed one and
one-half hours by mandatory holds while trains

passed through the area. Terminal countdown time
was only 12 minutes, 45 seconds. Launch, first stage
THOR performance, seporation, AGENA ignition, and
orbital injection were excellent. Damage to the
launch pad was unusually fight.

@ THOR mainstage ond vernier cutoff were very
near nominal. AGENA separation was clean and pos-
itive. After the proper delay, AGENA engine ignition
occurred as planned. Data confirmed that the AGENA
engine was shut down by the integrator-accelerom-
‘eter at a total velocity gain very close to nominal. The
resulting orbit has a perigee of 109.5 statute miles,
an apogee of 380 statute miles, an eccentricity of
033, and an orbital period of 92.3 minutes. Acqui-
sition of the satellite on the first pass was within 2
seconds of the time predicted from ascent tracking.

@ Orbital performance was outstanding. Acquisl-
tion was accomplished by every station on every
pass. AH commands were received and verified (a
total of fifteen}. The horizon scanner, inertial refer-
ence package, and gas jet contro! system functioned
extremely well, resulting in excellent satellite attitude
stabilization. The satellite power supply, including
the two advanced design static inverters, performed
efficiently. The main batteries lasted through the
26th orbit.

© This was the most completely instrumenjed DIS-
COVERER flown to date. Telemetry transmission was.
excellent and ail stations acquired valuable data on
satellite equipment functions. Telemetered satellite
internal temperatures were well within specified lim-
its, confirming modifications made to satellite thermal
control design.

@ In addition to a completely equipped AGENA
‘and recovery capsule, an experiment was flown with
@' precision dual-frequency doppler beacon from
the TRANSIT program. The beacon was accompanied
by a programmed continuous-burning light visible at
night as a seventh-magnitude star. This light was

Figure 5. DISCOVERER XI just after liftof.

SSR" NARENRP


@ foreword fo...



‘WDLPM-4 200

MONTHLY PROGRESS
COMMUNICATIONS. SATELLITE Program

Program Administration

@ AFBMD is Preparing a new development plan in
accordance with the requirements defined in amend-
ment 4 (dated 29 February 1960) to ARPA Order
No. 54. It is anticipated that this plan will be sub-
mitted to ARPA early in May,

Technical Progress

@ As directed by amendment 4, @ pretiminary sys-
tem analysis and design study has been made as the
first step toward providing for the interchangeable
installation of UHF or microwave equipment in a
single final stage vehicle compatible with either
AGENA or CENTAUR stages. It is considered that, in
Gecordance with this requirement, a similar final
stage vehicle could be boosted into o é-hour polar
orbit using the ATLAS/AGENA “Br combination,
Providing the following modifications were made:

a. Redesign the aerodynamic fairing to provide
70-inch maximum diameter final stage vehicle.

b. Provide a minimum velocity increment capa-
bility of 1,200 ft/sec. for the final stage vehicle
following separation from the AGENA stage.

@ General Electric Missile and Space Vehicle
Department is continuing a minimum sustaining effort
Pending full epproval and funding of the ADVENT
Program. Preliminary design studies and laboratory
testing are being continued. In addition, radiation
testing of bearing lubricants has been initiated and
several preliminary test teports, specification and
related documents have been issued. Preliminary
design of subsystems is also being continued.

© A temperature analysis of the final stage vehicle
solor array reveals that temperature gradients in the

array caused by entry into and exit from an eclipse
are negligible.

© Results of an eclipse study reveal that in a 24-
hour equatorial orbit, the satellite is In the umbra
and penumbra a maximum 74 minutes. The eclipse
season is 48 hours long, followed by a period of 135
days during which there is no eclipse,

UHF Communications Equipment

@ Anew subsystem description has been Prepared
based on the concept of vaing a similar final stage
vehicle for both UHF and microwave communications
repeater equipment.

@ WADD/Bendix visited Sylvania Corporation on
15-16 March to observe and discuss Progress under
@ WADD contract covering development of ultra-
reliable micor-miniature receiver circuitry. Arrange-
ments were made for continuing exchange of
information between Bendix and Sylvania,

Microwave Communications Equipment

@ USASRDL personnel visited AFBMD on 7-8 March
to review and discuss the AFBMD Prepared develop-.
ment plan, Funding estimates for all Phases of the
Program were agreed upon. Technical data covering
the microwave communications, as submitted to
AFBMD by USASRDL, were included in the develop-
ment plan with only minor deviations which will be
corrected before the plan is submitted to ARPA.

e Preliminary test results obtained on the Varian
Associates 8,000 mc, high-power klystron indicate
that the tube is capable of the required performance.

© Draft specifications for the ground complex and
satellite equipment have been completed by
USASRDL.



Payloads may be Installed on the AGENA forward
equipment rack, or distributed throughout the vehicle,
An ejectable recovery capsule has been d

for the nose section of the AGENA vehicle and flight
tested in the DISCOVERER Program.

NOSE CONE

/—FUEL AND
OXIDIZER
TANK
SECTION .

R

AGENA “g"
1,600
13,106


“Scbead & re aad trans-
ommend receiver on upper right).

Pigure 8, Eo eped SAMOS
Eitri i i Et est

mine pt:

+

. Set rk pe

{menace eigen ana
nai i

v¢é
74



Activities summarized in this report include the major space systems,
Projects and studies for which the Air Force Ballistic Missle Division ic
J . wholly oF partially responsible. Each space system and project is precacied
i by a concise history of administration, concept and objectives, making the

WDLPM- 200

WDAI 60 2028

~ . - we



WDLPM-4 200

CONFIDENTIAL

Pre-Launch

@ = ABLE-4 THOR pre-launch activity began in July
1959 when THOR 219 was installed on launch ped

17A. The booster remained on the pad until it lifted |

off under its own power eight months later. This
Period encompassed the most extensive testing and
checkout program ever conducted on a single vehicle.

@ The second stage used for ABLE-4 THOR was
originally programmed for and installed on the
ABLE-4 ATLAS. This stage was damaged while being
removed from the ATLAS; repaired, and installed on
THOR 219, When the ABLE-4 THOR flight was
deferred by NASA, the second stage wal removed
and stored. With the resumption of pre-launch opera-
tions in February, this stage developed a leak which
was repaired after removal of the engine and six
hours of intensive X-raying. On 25 February, a thrust
chamber norzle closure diaphragm was blown acci-
dentally. The stage was again repaired and
reinstalled on THOR 219 on 29 February. No further
problems were encountered.

@ During inspection of the third stage prior to final
installation on the flight vehicle, two gouges were
discovered in a critical area of the casing. The back-
vp third stage for ABLE-3, stored for nearly a year in

Payload and Satellite
Vehicle Assembly . ..

Technicians installing components in payload platform
(above) and placing top shell of satellite vehicle

on payload platform (top right). Completed satellite
vehicle (bottom right) showing mounting clamp and
arms for solar paddle installation.

the AMR Ordnance area, was substituted. Evidence
of excessive third stage vibration obtained from
NASA SHOTPUT flights in December necessitated a
delay to permit investigation and corrective action.
Extensive studies of vibration, test procedures, and
test stand facilities were conducted. A revised testing
Program was formulated ond performed with satis-
factory results being obtained.

©@ Construction of the vehicle which eventually
became the PIONEER Y satellite was htarted over a
year ago, The finished test unit was shipped to the
AMR on 27 November 1959. As a result of compo-
nent failures in the power supply subsystem, it was
returned to Space Technology Laboratories (STL) on
19 December for evaluation and testing. During
January and February it was subjected to exhaustive
qualification testing both at STL and in an altitude
simulation chamber, The requalified vehicle arrived
at the AMR on 25 February.

@ On March 2, a power converter failed during
special calibration testing of the 5-watt transmitter
causing the flight to be rescheduled from 8 March to
10 March. On 10 March, difficulties with the LOX
supply topping equipment caused the countdown to
be stopped at T minus 5 minutes, and subsequent
rescheduling on the launch attempt for 11 March.



- would not delay or otherwise affect the DISCOVERER
flight test program. The first of these vehicles would
be launched into circular polar orbits from Vanden-

wen berg AFB around mid-September. No changes in

the DISCOVERER command circuitry or procedures
would be necessary. The AGENA vehicles would pro-
vide extended burn capability, and would orbit in a
nese-aft, horizontal position, with the scanner unit
mounted vertically. The payload would have a useful

. life of four days (eight passes per day). Each |, pass

- would provide an average payload readout time of
four minutes.

* Technical Progress
Second Stage Vehicle

© Manvfacture of the AGENA "B" vehicle for the
third MIDAS flight test has been started, with com-
pletion scheduled for 11 July. Launch is programmed
for December. A full scale mockup of the solar
auxiliary power array for this vehicle was completed
in March. Several functional tests of the mockup have
been completed successfully. Solar cell. manufacture
is in progress,

Infrared Scanner Units

Infrared scanner units for the first two MIDAS
_ fights are being manufactured by Aerojet
General Corporation, and for flights 3, 4 and 5
by Baird-Atomic, Inc.

Figare 8 Ihre rear No.3 (fore
jetGenerd) shown moweied

Misra) with 11
a Special cme ibe sige tide
seneperature simulatic brior to
testing.
WDLPM-4 200

© Thermal and altitude tests of the third Aerojet-
General Corporation infrared scanner unit were con-
ducted in the high altitude temperature simulation
chamber between 7 and 10 March. The scanner unit
‘was mounted in a specially designed collimator dur-
ing the tests. A black box source of infrared radiation
also was installed inside the chamber. Readout of
infrared signals was performed successfully at vari-
ous temperatures and pressures, and over the full
range of scanner head positions. Scanner unit Opera-
tion also was satisfactory throughout the tests.

@ Tests of the thermomechanical equivalent model
of the Baird-Atomic scanner unit were completed in
the high altitude temperature simulation chamber on
4 March. At low temperatures some binding in the
main bearing occurred. A low temperature testing
Program has been initiated as an effort toward solv-
Ing this problem.

© The first Baird-Atomic scanner unit is under-
going optical testing and the third unit is being
assembled.

Communications ond Control

© Qualification testing of the engineering test
model of the 256 channel (Mode! B} PAM multiplexer
is in progress. This. unit will be used on the third
MIDAS flight test. The multiplexer is essentially an
electronic commutator for the telemetry system, per-

Acer pc OT IN

c7



CONFIDENTIAL

imately. 48 seconds after launch. The ATLAS per-
formed as planned over its entire powered flight
trajectory, The trajectory of this flight, from the

Atlantic Missile Range to the vicinity of the moon,’

was established to achieve the tightest possible cir.
cular iynar orbit consistent with the highest proba-
bility of success, The final bumout conditions were
to have provided an inertial velocity of 34,552 feet
er second. The payload was designed to investigate
space environment and Propagation effects and to
transmit crude television images of the far side of
the moon. This was the frst fight in which an ATLAS
ICBM was wed as the booster for a mullistoge

ABLE-4 THOR is a deep space probe which was
launched into a solar orbit intermediate between the
orbits of earth ond Venus. At its closest approach
to the sun, the probe will pass near the orbit of
Venus, returning fo 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 fight. The payload contains no retro:
rocket, and no attempt will be made to intercept

fenus. The design of the Payload components is the
same as those which have been proven on the
ABLE-3 satellite,

es padi demwing of ABLE-4 ATLAS fight te while. Photos sow vebicle on
re subsequently

lennch pad. Vebicle booster
firing readiness,

CONFIDENTIAL

and second-stage were destroyed during Might

WDLPM-4 200


WDLPM.4 200

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, space has been divided into three broad
Greas; earth orbital, lunar, interplanetary. Studies

in the Lunar area: SR 192 {U) Strategic Lunar Sys-
fem, 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 o 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 alt facets of military opera-
tions such as offensive, defensive, and 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 surveitiance 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
entice earth and 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 ond their probable ranges
will be examined. This study will also include the
means of logistically supporting and establishing the
lunar bose. This study was funded with $420,000 in
Fiscal Year 1959. Three contractors were funded
and three additional contractors are performing the
study on a voluntary basis, Consequently, it is esti-
mated that this study has the equivalent of $1.5
millions being applied to it,

4. The interplanetory 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 funar area
because the operational problems involved appear
to be somewhat different, the distances are much
Greater; our present knowledge of the Grea 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-

tractors’ final reports are due at AFBMD in Feb-

tuary 1960.

Nl


ENTER ELLIPTICAL ORBIT
PERIGEE OF 7

°
COAST ELLIPSE Yo 180) NN __
- ~
¢ - ~
AGENA \ ~,
1ST. BURN:
\ 150 N. Mi.

VERNIER SEPARATION &

AGENA 2 ND. BURN
ENTER

=m = oo

¢
Figure 2. . attitude reference. Also governs velocity magnitude and
Launch-to-orbit trajectory for flights 3 and subsequent, Opti- direction by inertial guidance system monitored by borizon
mum ATLAS boost, spied by radio-inertial system. scanner. Orbital attitude maintained by reaction wheel and
AGENA ascent (coast, burn, coast, second burn) provides gas jets. .

-_

PROGRAM HISTORY

The MIDAS Program was included in Weapon
System 127L when WS 1171 was transferred
fo the Advanced Research Projects Agency
early in 1959. ARPA subsequently separated
WS 1171 into the DISCOVERER, SAMOS and
MIDAS Programs, with the MIDAS objectives
based on an infrared reconnaissance system.
The MIDAS (Missile Defense Alarm System) ©
Program was directed by ARPA Order No.
38, dated 5 November 1958 until transferred
fo the Air Force on 17 November 1959.
An ARDC development plan for a ten flight
R&D program has been approved. This R&D
program will make possible the achievement
of a reliable operational system by 1963.

Figore 3.

Proposed MIDAS system. Four satellites spaced equidistant in each of yy
seo orbogona lanes at 2,000 2.m. alsisude. Pe ‘Maximum COv- vi

erage of USSR bien minimum number of sate
; WDLPM-4 200

c2



DISTRIBUTION

Headquarters, United States Air Force. 26
Headquarters, Air Research and Development Command 10
Strategic Air Command i]
Air Force Cambridge Research Center . 1
Air Force Flight Test Center J
Rome Air Development Center 1
Air Force Missile Development Center 1
Wright Air Development Division 1
Air Force Special Weapons Center 1
Air University

Amold Engineering Development Ceriter
Air Proving Ground Center

Alr Defense Command

Air Training Command

Air Photo and Charting Service

Air Force Missile Test Center

United States Air Force Academy

Air Technical Intelligence Center

6555th Test Wing (Development)

Air Force Ballistic Missile Division (ARDC)
Ballistic Missiles Center (AMC)

Assistant CINCSAC (SAC MIKE)
‘Aeronautical Chart and information Center
Rand Corporation

San Bemardino Air Materiel Area

6594 Test Wing (Satellite)

1002 Insp. Gen, Group

3415 Technical Training Group

Technical Air Command

7S e-a-~-- owe

~~-- oe org

WDLPM4 200 SECRET ~ CONFIBEN Ft



detailed drawing of AG:

AGMA
SRCOND STAGE “an —
Weight inert 1370 1400
Impote Propalionts 6590 13.100
“Feet Wong
Onidicer {1EFNA)
Pyrotechnics C4 100.

GROSS WEIGHT (tbs) 707 14,800

Engine YURBIoS XUR81to-7
Tarest, tbe. 15,000 15,000
‘Spee. tmp., sec. fvac) a7 a7
Bern Time, sec. 120 40
Restart Provisions No Yes

THOR BoosTER S475 DM.21
Weight—Dry 6,980 5,950
Ful 33750 33750
Oxidizer (LON) 4,200 _ 68,200
GROSS WEIGHT (iu) “109,000 ~t08,000

MB-3 B-3

Engine Block! = Block 2
Thrust, os. (8.1) 182900 167,000
Spec. tmp,, see. (5.1) ure 8
Born Time, sec. 163 163

Figure 1. Photograph o; ‘wo-stage DISCOVERER vehicle (left) and

A, second stage (right).

GAS JETS (ROLL)
GIMBALLED ENGINE
VELOCITY METER
—
MMAND 5 ~~
\fiwe TO FIRE
COMMAND 6

VELOCITY TO
BE GAINED

~

/— FUEL AND
OXIDIZER

TANK
{ SECTION

NITROGEN
SPHERE

‘SEPARATION
PLANE

~) HELIUM
SPHERES

ULLAGE ROCKETS
ENGINE (UDMH)



‘WDLPM-4 200

Summary of... /
“FIRSTS in SPACE”

ABLE4 THOR
PIONEER V

1, The greatest range over which man has maintained command of a“

2.

3

6.

an instrumented space vehicle.
The Greatest range over which a man-made object has been “
tracked.

The first instrumented space laboratory which has Permitted meas-
urements to be made of interplanetary space beyond 250,000°

miles in the following areas: —

a. The number of meteorites.

b. The momentum spectrum of these Interplanetary dust partictes.

¢. The intensity of the interplanetary magnetic field.

d. The planar direction of this field.

@. The total energy of radiation in interplanetary space.

f. The two-level energy levels of this radiation.

. The first test of on interplanetary guidance system components

(with no force contro! of the satellite available),

. The first interplanetary space probe to carry its own, self-sustaining

auxiliary power supply.

The highest powered transmitter {150-watts) ever carried on a
space probe mission.

E25



WDLPM-4 200

MONTHLY PROGRESS—AGENA Program
AGENA “1g
XLR81-Ba-7 Engine

© The formal review on 15-17 March of the
Preliminary Flight Rating tests resulted in « require-
ment for additional firings to test effects of propellant
water content on engine performance and to demon-
strate satisfactory engine operation following the
malfunction condition of simultaneous firing of both
starting charges. These additional firings were suc-
cessfully completed during March. Formal inspection

Of the disassembled engines in April will conclude ~

the Preliminary Flight Rating tests of this engine.

XLR81-Ba-9 Engine

© Testing of the XLRB1-Ba-9 engine installed in
the propulsion vehicle test Ossembly was completed
at Santa Cruz Test Base. The fotal testing Program
consisted of nine firings, two of which were com-
pleted in February. All firings included gimballing
tests and propellant utilization evaluation, A weld
crack had developed in the thrust chamber during
the second gimbailing test on 25 February. New
welding techniques and more stringent quality con-
trol procedures will eliminate this problem. Five
subsequent firings were made during 2-10 March

‘duration used the repaired original thrust chamber
and were completed 14 March, Problems with thrust
chamber erosion were ‘encountered in the tests on
the replocement thrust chamber. Bell Aircraft Cor-
poration is conducting on intensive experimental
Program to determine the effectiveness of new thrust
chamber coatings and to determine Possible injector
modifications to relieve the thrust chamber cooling
problems. in the interim, however, adequate thrust
chamber cooling can be @ssured by closely contro}.
ling the solids content and the ambient temperature
Of the oxidizer used.

© Testing of the ‘engine uncooled nozzle extension
{to achieve a 45:1 area ratio) was Continued at Bell

fired last month at Bell was sent to AEDC and suc-
cessfully tested in three separate firings of 85 sec-
onds duration each during 21 - 25 March,

13


Installing Solar Cell Paddies

Third Stage...

Payload installed in
Helmholtz Coil Test fixture,
used to check operation
of magnetometer
experiment .

E12

..» Satellite
Vehicle

Checking Paddle
Latching Mechanism

Third stage and
Satellite Vehicle,
Paddles folded
Down. Half of
Fairing Shell in
Place.



Installation on Launch Pad

Second Stage...

Delivered to gantry {above}. . . being raised into
gantry tower (above left} . . . being lowered into mating
Position with THOR (bottom left)... and during

mating with THOR.

WDLPM-4 200


A. THOR—SM-75 / AGENA “A‘*

B. THOR—DM-21 / AGENA "BY C, THOR—DM-21 / AGENA “B’

MB-3 Block 1 / XLR81-Ba-7 MB-3 Block 2/ XLR81-Ba-9
Flight History:
OISCOVERER AGENA THOR
No. Ne. Ne. Flight Dote Remorks
° 1019 160 21 Ja AGENA disizo: function
. mary THON sorbed for uot ot Rgho
1 1022 163 28 Feb 1959 rad orbit success{ully, Telemetry received for 314 seconds after
H) 1018 170 13 April jed on 17th orbit
Avi Certara cats rm ce et
i) 1020 174 3s Lannch, fhe on opti orbital boost . Failed
we harap eee ae

Vv 1023 7g 25 June Same as DISCOVERER Ill.

v 1029 192 13 August ALL objectives successfully achieved except capsule recovery after
jection on 17tb orbit,

vt 1028 200 19 August Same as DISCOVERER V.

vil 1051 206 7 November Attaived orbit success Lach 400-63 revented
Viliaaion on orbit na a poe a 4 rele bower pr “

vil 1050 212 20 November —  Astained orbit , hi AGENA
std, Mas el GaN

ix 1052 218 4 February THOR shut dows prematurely. Umbilical cord mast did not retract.
Quick disconnect failed, canting loss of beliume pressure.

x * 1054 223 19 February THOR destroyed at T plus 56 s0c. by Range Safety Officer.

“<CONEDENTIAL — qQagaganeTTM

AS