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NASA SP-6

RESULTS OF THE SECOND
U.S. MANNED
ORBITAL SPACE FLIGHT
MAY 24, 1962

NATIONAL AERONAUTICS
AND SPACE ADMINISTRATION
MANNED SPACECRAFT CENTER
PROJECT MERCURY

For sale by the Superintendent of Documents, U.S, Government Printing Office
Washington 25, D.C. - Price 65 cents

FOREWORD

This document presents the results of the second United States manned
orbital space flight conducted on May 24, 1962. The performance discussions
of the spacecraft and launch systems, the modified Mercury Network, mission
support personnel, and the astronaut, together with analyses of observed
space phenomena and the medical aspects of the mission, form a continuation
of the information previously published for the first United States manned
orbital flight, conducted on February 20, 1962, and the two manned sub-
orbital space flights.

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CONTENTS

Page

FOREWORD - 111

1. SPACECRAFT AND LAUNCH-VEHICLE PERFORMANCE 1

By John H. Boynton, Mercury Project Office, NASA Manned Spacecraft
Center; and E. M. Fields, Mercury Project Office, NASA Manned Space-
craft Center.

2. MERCURY NETWORK PERFORMANCE 15

By James J. Donegan, Manned Space Flight Support Division, NASA Goddard

Space Flight Center; and James C. Jackson, Manned Space Flight Support
Division, NASA Goddard Space Flight Center.

3. MISSION OPERATIONS

By John D. Hodge, Asst. Chief for Flight Control, Flight Operations Division,
NASA Manned Spacecraft Center; Eugene F. Kranz, Flight Operations
Division, NASA Manned Spacecraft Center; and William C. Hayes, Flight
Operations Division, NASA Manned Spacecraft Center.

4. SPACE SCIENCE REPORT 35

By John A. O’Keefe, Ph. D., Asst. Chief, Theoretical Division, NASA Goddard
Space Flight Center; Winifred Sawtell Cameron, Theoretical Division, NASA
Goddard Spice Flight Center.

5. AEROMEDICAL STUDIES 43

A. CLINICAL MEDICAL OBSERVATIONS 43

By Howard A. Minners, M.D., Aerospace Medical Operations Office,

NASA Manned Spacecraft Center; Stanley C. White, M.D., Chief,

Life Systems Division, NASA Manned Spacecraft Center; William K.
Douglas, M.D., Air Force Missile Test Center, Patrick Air Force Base,

Fla.; Edward C. Knoblock, Ph. D., Walter Reed Army Institute of
Research, Washington, D.C.; and Ashton Graybiel, M.D., U.S. Naval
School of Aviation Medicine, Pensacola, Fla.

B. PHYSIOLOGICAL RESPONSES OF THE ASTRONAUT .54

By Ernest P. McCutcheon, M.D., Aerospace Medical Operations Office,

NASA Manned Spacecraft Center; Charles A. Berry, M.D., Chief,
Aerospace Medical Operations Office, NASA Manned Spacecraft
Center; G. Fred Kelly, M.D., U.S. Naval Air Station, Cecil Field,
Jacksonville, Fla.; Rita M. Rapp, Life Systems Division, NASA
Manned Spacecraft Center; and Robie Ilackworth, Aerospace Medical
Operations Office, NASA Manned Spacecraft Center.

6. PILOT PERFORMANCE 63

By Helmut A. Kuehuel, Flight Crew Operations Division, NASA Manned
Spacecraft Center; William O. Armstrong, Flight Crew' Operations
Division, NASA Manned Spacecraft Center; John J. Van Iiockel,

Flight Crew Operations Division, NASA Manned Spacecraft Center;
and Harold I. Johnson, Flight Crew Operations Division, NASA

Manned Spacecraft Center.

7. PILOT'S FLIGHT REPORT 69

By M. Scott Carpenter, Astronaut, NASA Manned Spacecraft Center.

APPENDIX. MA-7 AIR-GROUND VOICE COMMUNICATION .. 77

1. SPACECRAFT AND LAUNCH-VEHICLE PERFORMANCE

By John H. Boynton, Mercury Project Office, NASA Manned Spacecraft Center; and E. M. Fields,
Mercury Project Office NASA Manned Spacecraft Center

Summary

The performance of the Mercury spacecraft
and Atlas launch vehicle for the orbital flight
of Astronaut M. Scott Carpenter was excellent
in nearly every respect. All primary mission
objectives were achieved. The single mission-
critical malfunction which occurred involved a
failure in the spacecraft pitch horizon scanner,
a component of the automatic control system.
This anomaly was adequately compensated for
by the pilot in subsequent inflight operations so
that the success of the mission was not- compro-
mised. A modification of the spacecraft con-
trol-system thrust units were effective. Cabin
and pressure-suit temperatures were high but
not intolerable. Some uncertainties in the data
telemetered from the bioinstrumentation pre-
vailed at times during the flight; however, as-
sociated information was available which indi-
cated continued well-being of the astronaut.
Equipment was included in the spacecraft
which provided valuable scientific information;
notably that regarding liquid behavior in a
weightless state, identification of the airglow
layer observed by Astronaut Glenn, and photo-
graphy of terrestrial features and meteorolo-
gical phenomena. An experiment which was
to provide atmospheric drag and color visibility
data in space through deployment of an inflata-
ble sphere was partially successful. The flight
further qualified the Mercury spacecraft sys-
tems for manned orbital operations and pro-
vided evidence for progressing into missions of
extended duration and consequently more de-
manding systems requirements.

Introduction

The seventh Mercury- Atlas mission (MA-T)
was planned for three orbital passes and was a
continuation of a program to acquire operation-
al experience and information for manned or-

bital space flight. The objectives of the flight
were to evaluate the performance of the
man-spacecraft system in a three-pass mission,
to evaluate the effects of space flight on the
astronaut, to obtain the astronaut’s opinions on
the operational suitability of the spacecraft sys-
tems, to evaluate the performance of spacecraft
systems modified as a result of unsatisfactory
performance during previous missions, and to
exercise and evaluate further the performance
of the Mercury Worldwide Network.

The Aurora 7 spacecraft and Atlas launch
vehicle used by Astronaut Carpenter in success-
fully performing the second United States
manned orbital mission (MA-7) were nearly
identical to those used for the MA-6 flight. The
Mercury spacecraft provided a safe and habita-
ble environment for the pilot while in orbit, as
well as protection during the critical flight
phases of launch and reentry. The spacecraft
also served as an orbiting laboratory where the
pilot could conduct limited experiments which
would increase the knowledge in the space sci-
ences. The intent of this paper is to describe
briefly the MA-7 spacecraft and launch vehicle

Figure 1-1. — Mercury spacecraft systems.

systems and discuss their technical perfor-
mance.

The many systems which the spacecraft com-
prises may be generally grouped into those of
heat protection, mechanical and pyrotechnic,
attitude control, communications, electrical and
sequential, life support, and instrumentation.
The general arrangement of the spacecraft in-
terior is schematically depicted in figure 1-1.
Although a very basic description of each sys-
tem accompanies the corresponding section, a
more detailed description is presented in refer-
ence 1.

Heat Protection System

During flight through the atmosphere at
launch and reentry, the high velocities generate
excessive heat from which the crew and equip-
ment must be protected. The spacecraft must
also be capable of withstanding the heat pulse
associated with the ignition of the launch es-
cape rocket. To provide this protection, the
spacecraft afterbody is composed of a double-
wall structure with thermal insulation between
the two walls. The outer conical surface of the
spacecraft afterbody is made up of higli-tem-
perature alloy shingles, and the cylindrical por-
tion is protected by beryllium shingles. The
spacecraft blunt end is fitted with an ablation-
type heat shield, which is constructed of glass
fibers and resin. Additional description of the
heat protection system can be found on pages
7 to 9 of reference 1.

Although the MA-7 reentry trajectory was
slightly more shallow than for MA-fi, the heat-
ing effects were not measurably different, as is
evident in figure 1-2.

The performance of the MA-7 heat protec-
tion system was as expected and was quite sat-
isfactory.

2,000

1,500

° 1,000
&

500 h

— Predicted

O MA-4

□ MA-5

o MA-6

A MA- 7

Inside

surface-

- Outside
surface

"Ml — & ~ l

40 60

TTuckness, percent

100

FniritE 1-2. — Maximum ablation-shield teniperul tires.

Two temperature measurements were made
in the ablation shield, one at the center and the
other approximately 27 inches from the center.
The maximum recorded values are graphically
shown and compared with previously obtained
orbital reentry values in figure 1-2. The mag-
nitudes of these temperatures, as well as the
ablation-shield weight loss during reentry, are
comparable with previous flights. The sup-
porting structure behind the ablation shield was
found to be in excellent condition following the
flight. A more complete temperature survey
of points around the afterbody than on previ-
ous flights was conducted for MA-7. This sur-
vey was made possible by the addition of a low-
level commutator circuit. The data, which are
shown in figure 1-3, were within expected
ranges, and the integrity of the structure was
not affected by the thermal loads experienced.

Mechanical and Pyrotechnic Systems

The mechanical and pyrotechnic system
group consists of the separation devices, the
rocket motors, the landing system, and the in-
ternal spacecraft structure. This entire group
functioned satisfactorily during the mission.
Performances of individual systems are dis-
cussed in the following paragraphs.

Separation Devices

Separation devices generally use explosive
charges to cause separation or disconnection of
components. The major separation points are
at the interface between the spacecraft and
launch vehicle, between the spacecraft and the
escape tower, at the heat shield, and around
the spacecraft hatch. All separation devices
worked properly during the mission. The ex-
plosive-actuated hatch was not used, since the
pilot egressed through the top of the spacecraft
after landing,

Rockei Motors

The rocket motor system consisted of three
retrorocket motors, three posi grade motors, the
launch escape rocket, and the small tower- jetti-
son rocket. All of these motors are solid-pro-
pellant type. See page 9 of reference 1 for
additional description of the rocket motor sys-
tem. Nominal thrust and burning-time data
are given in the following table :

Although ignition of the retrorocket motors
was about 3 seconds later than expected, the per-

Pocket motor

Number

motors

Nominal
thrust
each, lb

Approx-

imate

burning

time

each,

sec

Escape - _ -

52, 000

Tower jettison.. . .

800

1. 5

Posigrade..

400

Retrograde...

1, 000

formance of the rocket motors was satisfactory.
An analysis of radar tracking data for the
flight, yielded a velocity increment at retrofire
which indicated that the retrorocket perform-
ance was 3 percent lower than nominal. This
was acceptable and within the allowable devia-
tion from nominal performance of ±5 percent.

Landing System

The landing system includes the drogue
(stabilization) parachute, the main and reserve
parachutes, and the landing shock-attenuation
system (landing bag). The latter system at-
tenuates the force of landing by providing a
cushion of air through the deployment of the
landing bag and heat shield structure, which is
supported by straps and cables. The landing
system can be actuated automatically, or manu-

ally by the astronaut. The landing system is
described in greater detail on pages 28 to 30 of
reference 1. The landing system performed
satisfactorily and as planned.

The MA-7 landing system differed from the
MA-6 system in the manner of arming the
barostats (pressure -sensing devices). These
units initiate automatic deployment of the para-
chutes when the spacecraft descends to the
proper pressure altitude during reentry. In
the MA-6 and prior missions, the barostats
were armed when above the atmosphere dur-
ing exit flight and thus were in readiness to
initiate the parachute-deployment mechanisms
when the barostats sensed the appropriate pres-
sure during spacecraft descent through the at-
mosphere. This armed status of the barostats
would of course permit deployment of a para-
chute during orbital flight if a certain type of
barostat malfunction should occur. While
such barostat malfunctions had not been de-
tected in previous flights or during ground
tests, it was believed that an additional safety
margin would lie desirable, because of the un-
explained early deployment of the drogue para-
chute during the MA-6 mission. Consequently,
a control barostat was added to the automatic
sequence circuit of the MA-7 spacecraft. This
barostat sensed pressure in the cabin and func-

Figtjre 1-3. — Afterlxxly temperature from low-level commutator circuit.

tioned in a manner intended to prevent auto-
matic deployment of the parachutes until the
spacecraft cabin pressure corresponded to an ac-
ceptable altitude. The control barostat did not
alter the circuitry that was available to the pi-
lot for manual deployment of the parachutes.
In the MA T mission, it was planned for the pi-
lot to deploy the drogue parachute manually
at an altitude, of 21,000 feet or higher if added
spacecraft stabilization prior to automatic de-
ployment of the drogue parachute was desired.
Astronaut Carpenter felt the need for addi-
tional spacecraft stabilization . and manually
deployed the drogue parachute at an altitude of
approximately 25,500 feet. The pilot also
planned a manual deployment of the main
parachute routinely at an altitude of about
10,000 feet, rather than waiting for automatic
deployment at approximately 8,200 feet alti-
tude; the data show that manual deployment
was effected at about 9,500 feet.

Flotation

After landing, the astronaut reported a severe
list angle on the order of 60° from vertical,
and postfligbt photographs of the spacecraft
taken after egress of the pilot indicate approxi-
mately a 45° list angle. The time normally re-
quired for the spacecraft to erect to its equilib-
rium angle exceeds the period that Astronaut
Carpenter used to initiate egress ; therefore, this
egress activity may have prevented the return
to a more nearly vertical flotation attitude.
Upon recovery, a considerable amount of sea
water was found in the spacecraft, the majority
of which is believed to have entered through
the small pressure bulkhead when the pilot
passed through the recovery compartment into
the liferaft. In addition, small leaks in the
internal pressure vessel which probably oc-
curred upon landing were disclosed during the
normal postflight inspection; but accounting
for the 6 hours prior to spacecraft recovery,
these leaks would have contributed only slightly
to the water in the cabin. The pilot reported
that at landing a small amount of water
splashed onto the tape recorder in the cabin; it
is believed that this resulted from a surge of
water which momentarily opened a spring-
loaded pressure relief valve in the top of the
cabin.

Spacecraft Control System

The spacecraft control system is designed to
provide attitude and rate control of the Mer-
cury vehicle while in orbit and during reentry.
Page 11 of reference 1 presents an additional
description of the spacecraft control system.
With the single exception of the pitch horizon
scanner, all system components performed nor-
mally during the entire flight.

Table 1 - 1 . — Spacecraft Control System- Redun-
dancy and Electrical Power Requirements

Control

system

modes

Corresponding fuel
system (fuel sup-
ply, plumbing, and
thrusters)

Electrical

power

required

ASCS

d-c and a-c
d-c
None

d-c and a-c

FBW

A -

B - - -

RSCS

ASPS— Automatic stabilization and control system

FBW — Fly-by-wire 1

MP — Manual proportional Controlled by pilot

system ( actuation of eon-

RSCS — Rate stabilization con- trol stick
trol system /

The attitude control system, at the discre-
tion of the pilot, is capable of operation in the
modes listed in table 1-1.

The spacecraft was equipped with two sepa-
rate reaction control systems (RCS) shown as
A and B in table 1— I, each with its own fuel
supply and each independent of the other.
Combinations of modes ASCS and FBW, FBW
and MP, or FBW and RSCS were available to
provide “double authority” at the choice of the
pilot. A “maneuver” switch was added to the
instrument panel for MA-7 and was included
in the control circuitry to allow the astronaut
to perform maneuvers without introducing er-
rors in his attitude displays. Actuation of the
switch effectively disabled the yaw reference
slaving system and the automatic pitch orbital
precession of 4°/ min and thus prevented gen-
eration of erroneous gyro slaving signals dur-
ing maneuvers.

The reaction control components were of the
standard configuration, with the exception of
the 1-pound and 6-pound thruster assemblies
which had been slightly modified to correct de-
ficiencies which occurred on earlier flights. The

Modified

Figure 1-4. — Comparison between MA-6 and MA-7
thrusters.

modification to the 1-pound units involved re-
placing the stainless-steel fuel distribution
(Dutch weave) screens (see fig. 1-4) with plat-
inum screens and a stainless-steel fuel distribu-
tion plate, reducing the volume of the heat bar-
riers of the automatic RCS, and moving the
fuel-metering orifice to the solenoid inlet. The
only modification to the 6-pound units was the

replacement of the stainless-steel screens with
platinum screens. These changes proved to be
effective, as all thrust units operated properly
throughout the flight.

Horizon Scanners

The horizon scanners are employed to pro-
vide a correction reference for the spacecraft
attitude gyros which is indicated in the basic
schematic diagram shown in figure 1-5. An
error introduced by the pitch horizon scanner
circuit was present during launch and apparent-
ly remained to some degree throughout the
flight. Since the scanners were lost when the
antenna canister was jettisoned during the nor-
mal landing sequence, postflight inspection and
analysis of these units were impossible; how-
ever, the failure is believed to have been in the
scanner circuit and was apparently of a random
nature in view of the fact that the scanner sys-
tem has been fully qualified on previous flights.

Figure 1-5. — Control system schematic diagram.

Elapsed time, mTnsec

Figure 1-6. — Spacecraft attitudes during launch.

Some 40 seconds after escape tower separa-
tion, the output of the pitch scanner indicated a
spacecraft attitude of approximately 17°, which
is graphically depicted in figure 1-0. Also
shown is the attitude of the launch vehicle and
spacecraft as determined from launch vehicle
data which is about. — 1° at this time, indicating
a scanner error of about 18° in pitch. This er-
ror apparently increased to about 20° at space-
craft separation. Radar tracking data at the
time of retrofire provided the only additional
independent information source and the radar
data verify, in general, the scanner error. The
thrust vector which produced the postretro-
grade velocity was calculated by using the radar
measurements, and since this vector is alined
with the spacecraft longitudinal axis, a retro-
fire attitude of about —36° in pitch was de-
rived. This calculated attitude was compared
with a scanner-indicated attitude of —16° dur-
ing retrofire, yielding a difference of 20°. Al-
though these two independent measurements
and calculations would support a theory that
a constant bias of about 20° was present, the
attitudes as indicated by instruments and com-
pared with observations by the pilot disclose
possible horizon scanner errors of widely vary-
ing amounts during the orbital flight phase.
Because of the malfunctioning scanner which

resulted in pitch errors in the spacecraft atti-
tude-gyro system, the pilot was required to as r
sume manual control of the spacecraft during
the retrofire period.

Fuel Usage

Double authority control was inadvertently
employed at times during the flight, and the
fly-by-wire high thrust units were accidently
actuated during certain maneuvers, both of
which contributed to the high usage rate of
spacecraft fuel indicated in figure 1-7. In
addition, operation of the ASCS mode while
outside the required attitude limits resulted in
unnecessary use of the high thrust units. The
manual-system fuel was depleted at about the
end of the retrofire maneuver, and the auto-
matic-system fuel was depleted at about half-
way through the reentry period.

Because of the early depletion of automatic-
system fuel, attitude control during reentry
was not available for the required duration.
As a result, attitude rates built up after the
ASCS became inoperative because of the lack
of fuel, and these rates were not sufficiently
damped, as expected, by aerodynamic forces.
These oscillations to diverge until the pilot
chose to deploy the drogue parachute manually
at an altitude of approximately 25,000 feet to
stabilize the spacecraft.

In order to prevent inadvertent use of the
high-thrust jets when using FBW mode of con-
trol, the MA-8 and subsequent spacecrafts will
contain a switch which will allow the pilot to
disable and reactivate the high-thrust units at
his discretion. An automatic override will re-
activate these thrusters just prior to retrofire.
Additionally, a revision of fuel management
and control training procedures has been in-
stituted for the next mission.

Communication Systems

The MA-7 spacecraft communication system
was identical to that contained, in the MA-6
configuration with one minor exception. The
voice power switch was modified to provide a
mode whereby the astronaut could record voice
on the onboard tape recorder without trans-
mission to ground stations. Switching to the
transmitting mode could he accomplished with-
out the normal warmup time, since the trans-
mitter was maintained in a standby condition

when the switch was in the record position.
The communications system is described in
more detail beginning on page 12 of reference
1. The MA-7 communication system, with cer-
tain exceptions discussed below, performed
satisfactorily.

Voice Communications

The UIIF voice communications with the
spacecraft were satisfactory. Reception of IIF
voice in the spacecraft was satisfactory; how-
ever, attempts on the part of the astronaut to
accomplish ITF voice transmission to the
ground were unsuccessful. The reason for the
poor IIF transmission has not been determined.

Radar Beacons

Performance of the C- and S-band beacons
was entirely satisfactory, although slightly in-
ferior to that of the MA-6 mission. Several
stations reported some beacon countdown
(missed pulses) and slight amplitude modu-
lation on the C-band beacon. The amplitude
modulation was possibly caused by the modula-
tion presented by the phase shifter and the
drifting mode of the spacecraft, which resulted
in a less than optimum antenna orientation, as
expected. Both beacons were rechecked after

the mission and found to be essentially un-
changed from their preflight status.

Location Aids

The recovery beacons employed as postland-
ing location aids include the SEASAVE (HF/
DF), SARAH (UHF/DF), and Super
SARAH (UHF/DF) units. Recovery forces
reported that the auxiliary' beacon (Super
SARAH) and UHF/DF signals were received.
The Super SARAH beacon was received at a
range of approximately 250 miles. Both the
SARAH beacon and UIIF/DF transmitter
were received at ranges of 50 miles from the
spacecraft.

The SEASAVE rescue beacon (HF/DF)
was apparently not received by r the recovery
stations. The whip antenna used by this bea-
con was reported by the recovery forces to be
fully extended and normal in appearance. The
beacon was tested after the flight and found to
be operating satisfactorily. The reason for
lack of reception of this beacon has not been
established, but the large list angles of the
spacecraft after landing placed the whip an-
tenna near the surface of the water, and this
may have been a contributing factor.

Command Receivers

The two command receivers operated effec-
tively during the MA-7 flight. One exercise
was successfully accomplished with the emer-
gency-voice-mode of the command system while
over Muchea. The second exercise of this
mode, attempted during reentry, was unsuc-
cessful because the spacecraft was below the
line-of-sight of the range stations at this time.

Instrumentation System

The spacecraft instrumentation system was
basically the same as that for the MA-6 mis-
sion which is described on page 19 of reference
1. Performance of the system was satisfactory
except for those items discussed below.

The instrumentation system sensed informa-
tion pertinent to over 100 items throughout the
spacecraft. The biological parameters of the
pilot, namely electrocardiogram (ECG) traces,
respiration rate and depth, body temperature,
and blood pressure, were of primary concern
to flight control personnel. In addition, many
operational aspects of spacecraft systems were
monitored. These aspects included significant
sequential events, control system operation and
component outputs, attitudes and attitude rates,
electrical parameters, ECS pressures and tem-
peratures, accelerations along all three axes,
and temperatures of systems and structure
throughout the spacecraft. These quantities
were transmitted to Mercury Network stations
and recorded onboard the spacecraft. The
system also included a 16-mm motion picture
camera which photographed the astronaut and
surrounding portions of the spacecraft. The
instrumentation system had direct readouts on
the MA-7 spacecraft display panel for many of
the instrumented parameters.

System Modifications

The changes made since the MA-6 flight in-
cluded the incorporation of an additional, low-
level commutator circuit which provided a more
complete temperature survey, rewiring of the
circuitry which monitored closure of the limit
switches that sensed heat-shield release and the
substitution of a semi-automatic blood pressure
measuring system (BPMS) for the manual de-
vice used for MA-G. In addition, the earth-
path indicator and the instrument-panel cam-
era were deleted for MA-7.

Instrumentation Anomalies

A problem in the instrumentation system
occurred just after lift-off when erroneous ECG
signals were temporarily recorded. These ex-
traneous signals were primarily attributed to
rapid body movements of the pilot and possi-
bly excessive perspiration during this period.

For a short period during the orbital phase,
the instrumentation indicated that the astro-
naut’s temperature had increased to 102° F. and
this caused mojnentary concern. However,
other medical information indicated that this
102° F. value was erroneous. The respiration
rate sensor provided adequate preflight data,
but the inflight measurements were marginal
because of the variations in head position and
air density. This anomaly has been experi-
enced on previous flights and was of little con-
cern.

The data transmitted from the blood-pressure
measuring system were questionable at times
during the flight, primarily because of the mag-
nitude of the data and the intermittency with
which it was received. The intermittent sig-
nals were found to have resulted from a broken
cable in the microphone pickup, shown in figure
1-8. This malfunction, however, could not
have affected the magnitude of the transmitted
information, since an intermittent short either
sends valid signals or none at all. The BPMS
was thoroughly checked during postflight tests
in the laboratory using actual flight hardware,
with the exception of the microphone and cuff.
Tests of the controller unit and amplifier were
also conducted, and no component failure or
damage in the BPMS has been detected to date.
However, a number of uncertainties regarding
the calibration and operation of the BPMS still

Dump

solenoid

Occluding

: cuff ... .

1 t , Microphone

*' Suit Mercury

fitting .battery

Programed
: regulator
Fill

.solenoid
\°Z

MOOpsl

To spacecraft
telemetry

Preamplifier

Filters'

Figure 1-8. — Semiautomatic blood-pressure
measuring system.

exist. Extensive testing is being conducted to
correlate postflight and inflight BPMS read-
ings more accurately with clinically measured
values.

The remainder of the instrumentation sys-
tem performed satisfactorily, with the excep-
tion of a noncritical failure of one temperature
pickup, a thermocouple located at the low clock-
wise automatic thruster. A brief indication
of spacecraft descent occurred on the rate-of-
descent indicator toward the end of the orbital
phase ; but since this unit is activated by atmos-
pheric pressure, the indication was obviously
false. This indicator apparently operated sat-
isfactorily during descent through the atmos-
phere and was found to be operating properly
during postflight evaluation. The pilot-obser-
ver camera film suffered sea-water immersion
after the flight, and its quality and usefulness
were somewhat limited.

Life Support System

The life support system primarily controls
the environment in which the astronaut oper-
ates, both in the spacecraft cabin and in the
pressure suit. Total pressure, gaseous com-
position, and temperature are maintained at
acceptable levels, oxygen is supplied to the pilot
on demand, and water and food are available.
Both the cabin and suit environmental systems
operate automatically and simultaneously from
common oxygen, coolant water, and electrical
supplies. In-flight adjustment of the cooling
system is provided for, and the automatic-sup-
ply function of the oxygen system has a man-
ual override feature in case of a malfunction.

The suit and cabin pressures are maintained
at 5.1 psia, and the atmosphere is nearly 100-
percent oxygen. The environmental control
system (ECS) installed in the MA-7 space-
craft, schematically shown in figure 1-9, dif-
fered from that for MA-6 in only two respects :
The constant bleed of oxygen into the suit cir-
cuit was eliminated, and the oxygen partial
pressure of the cabin atmosphere was measured,
rather than that of the suit circuit. Pages 21
and 31 of reference 1 contain additional descrip-
tion of the life support system.

Higher-than-desired temperatures in the
spacecraft cabin and pressure suit were experi-
enced during the MA-7 flight, and these values

Figure 1-9. — Schematic diagram of the environmental
control system.

are plotted in figures 1-10 and 1-11, respec-
tively. In the same figures, the cabin and suit
temperatures measured during the MA-6 mis-
sion are shown for comparison. The high
temperatures were the only anomalies evident
in the ECS during the flight.

The high cabin temperature is attributed to

a number of factors, such as the difficulty of
achieving high air-flow rates and good circu-
lation of air in the cabin and vulnerability of
the heat exchanger to freezing-blockage when
high rates of water flow are used. Tests are
currently being made to determine if the cabin
temperature can be lowered significantly with-
out requiring substantial redesign of the cabin-
cooling system.

In the case of the high suit temperatures,
some difficulty was experienced in obtaining the
proper valve setting for the suit-inlet tempera-
ture control, mainly because of the inherent
lag at the temperature monitoring point with
control manipulations. The comfort control
valve settings are presented in figure 1-11, and
a diagonal line reflects a lack of knowledge as
to when the control setting was instituted. It
has been further ascertained in postflight test-
ing in the altitude chamber that the suit tem-
perature did respond to control valve changes.
Based on the satisfactory performance of the
suit system in the MA-G flight, it is believed
that the suit-inlet temperature could have been
maintained in the GO 0 to G5° F range for the
MA-7 flight, had not the comfort control valve

been turn'ed down early in the flight. The valve
setting was reduced by the pilot during the first
orbital pass when the cabin heat exchanger in-
dicated possible freezing. It is believed that
some freezing at the heat exchanger did occur
during the flight which may have resulted in
less efficient cooling, but this is not the primary
cause of the above-normal temperature.

A study of the metabolic rate associated with
a manned orbital flight was conducted in this
mission, and the results yielded a metabolic ox-
ygen consumption of 0.0722 pound/hour or 408
standard cubic centimeters (see) per minute.
This level under weightlessness is comparable
to that in a norma] gravity field with similar
work loads and is within the design specifica-
tion of 500 sec/min for the ECS.

Electrical and Sequential System

The electrical power system for MA-7 was
of the same type as that used for MA-G. This
system is described more fully on page 21 of
reference 1. The MA-7 electrical power sys-
tem performed satisfactorily during the MA-7
mission.

The sequential system for ]\f A-7 deviated only
slightly from that used for MA-6, which is
described in detail on page 26 of reference 1.
The major change involved the addition of a
control barostat in the landing sequence cir-
cuit, which is discussed in a previous section.
The MA-7 sequential system performed ade-
quately during the mission. The one anomaly
that was experienced is discussed subsequently.

Inverters

Temperatures of the inverters were, as in
previous flights, above expected values. How-
ever, a change in the coolant valve setting by
the astronaut later in the flight did decrease
the rate of rise in the inverter temperatures.

Squib Fuses

As expected, squib-circuit fuses were found
to be blown, including the number 1 retrorocket
switch fuse which also had a small hole on the
side of the ceramic housing. Postflight testing
demonstrated that at the electric current levels
experienced in flight, the casing of these fuses
could be ruptured and significant quantities of
smoke could be produced. It was confirmed
by the astronaut during postflight tests, where
lie observed two similar fuses being blown, that
these fuses produce a smoke having the same
color and odor as that encountered in flight at
the time of retrofire.

Sequential System

The differences between the MA-6 and MA-7
sequential systems included changing of the
horizon scanner slaving signal from programed
to continuous and locking-in of the %-g relay
at sustainer engine cutoff to prevent reopening
by posigrade thrust.

One sequential system anomaly was indi-
cated in the mission when retrofire was reported
to have been delayed about 1 second after the
pilot actuated the manual switch to ignite the
retrorockets. Figure 1-12 shows schematically
the retrosequenee circuitry. Since the attitude
gyro in pitch indicated that the spacecraft
pitch att itude was not within ±12° of the nom-
inal — 34°, the attitude-permission circuitry
could not pass the retrofire signal from the
clock and thus, the automatic clock sequence
could not ignite the retrorockets; this lack of
permission was proper and indicated that se-

quential circuitry performance was according
to design. After waiting for about 2 seconds,
Astronaut Carpenter actuated the manual retro-
fire switch. ITe reported that an additional
delay of about 1 second occurred before the
retrorockets actually ignited, which normally
would take place instantaneously. No explana-
tion is available for this additional 1-second
delay, since exhaustive postflight testing lias
failed to reveal any trouble source in the igni-
tion sequence circuitry.

Scientific Experiments

It was planned that a series of research ex-
periments would be conducted by Astronaut
Carpenter during the MA-7 mission. This
series included a balloon experiment, a zero-
gravity study, a number of photographic exer-
cises, a ground flare visibility experiment, and
observations of the airglow layer witnessed by
Astronaut Glenn. Results of the last ex-
periment are presented in paper 4 and will not
be discussed here. Most Mercury experiments
were proposed and sponsored by agencies out-
side the Manned Spacecraft Center. Each was
carefully evaluated prior to its approval for
inclusion in the flight plan. Sponsoring agen-
cies for the MA-7 experiments are shown in the
following table.

Experiment

Sponsoring organization

Balloon

Langley Research
Center

Zero-gravity _ _

Lewis Research Center

Ground flare visibility--

Manned Spacecraft
Center

Horizon definition

MIT Instrumentation
Laboratory

Meteorological

photography

U.S. Weather Bureau

Airglow layer .-

Goddard Space Flight
Center

Balloon Experiment

The objectives of the balloon experiment were
to measure the drag and to provide visibility
data regarding an object of known size and
shape in orbital space. The balloon was 30
inches in diameter, and was constructed of five
equal-sized lunes of selected colors and surface
finishes. The sphere was constructed of a plas-
tic and aluminum foil sandwich material, and

654533 0—62 2

'‘rfl?rosequ*nc«
manual control "

Figure 1-12. — Retrosequenee schematic diagram.

was to be inflated with a small nitrogen bottle
immediately after release from the antenna
canister at the end of the first orbital pass.
In addition, numerous hi -inch discs of alumin-
ized plastic were placed in the folds of the bal-
loon and dispersed when the balloon was de-
ployed. As intended, the pilot observed the
rate of dispersion and the associated visual ef-
fects of the “confetti.”

The balloon was deployed at 01 :38 :00 ground
elapsed time, but it failed to inflate properly.
The cause has been attributed to a ruptured
seam in the skin. Aerodynamic measurements
were taken with the strain-gage pickup, but
these are of little use since the actual frontal
area of the partial inflated balloon is not known.
The visibility portion of the experiment was also
only partially successful because only two of
the surface colors were visible, the orange and
aluminum segments. While the balloon was
deployed, a series of spacecraft maneuvers evi-
dently fouled the tethering line on the destabi-

lizing flap located on the end of the cylindrical
portion of the spacecraft, thus preventing the
jettisoning of the balloon. No difficulty was
encountered during retrofire and the balloon
burned up during reentry.

Zero-Gravity Experiment

The objective of the zero-gravity experiment
was to examine the behavior of a liquid of
known properties in a weightless state using a
particular container configuration. The ap-
paratus consisted of a glass sphere containing
a capillary tube which extended from the in-
terior surface to just past the center, as shown
in figure 1-13. A liquid mixture representing
the viscosity and surface tension of hydrogen
peroxide was composed of distilled water, green
dye, aerosol solution and silicon, and consumed
about 20 percent of the interal volume. The
diameter of the sphere was 3% inches. The
application of the results of this experiment is
primarily in the design of fuel tanks for fu-

Standpipe

Front view Side view Front view

(I-?) (I -g) (zero-?)

Figure 1-13. — Zero-gravity experiment.

ture spacecraft. The surface configuration of
the liquid under zero-gravity was expected to
assume the position indicated in the final view
of figure 1-13. An astronaut report during
the third pass over Cape Canaveral (see ap-
pendix) and a postflight analysis of the pilot-
observer film verified the predicted behavior of
the liquid. The results of the experiment
showed that the liquid filled the capillary tube
during weightless flight and during the low-ac-
celeration portion of reentry.

Ground-Flare-Visibility Experiment

The major objective of the ground-flare-
visibility experiment was to determine the ca-
pability of the astronaut to acquire and observe
a ground-based light of known intensity and
to determine the attenuation of this light source
through the atmosphere. The earth-based ap-
paratus consisted of ten 1,000,000-candle-power
flares located at Woomera, Australia. The pilot
was supplied with an extinction photometer
with a filter variation from 0.1 neutral density
to 3.8 neutral density (99.98-percent light reduc-
tion) . The flares, with a burning t ime of about
li/ 2 minutes, were to be ignited approximately
60 seconds apart during passes over this station.
The experiment was attempted and failed to
yield results because of heavy cloud cover during
the first pass. It was therefore discontinued for
the remainder of the flight because of continu-
ing cloud cover. This cloud cover, which was
also experienced during a similar experiment
in MA-6, was approximately eight-tenths at
3,000 feet. The exercise is scheduled to be re-
peated in a future flight.

Photographic Studies

A series of photographic exercises were
planned for the MA-7 flight, but since opera-
tional requirements assume priority over sched-

uled flight activities, some of these studies were
not conducted. The Massachusetts Institute of
Technolog}’ sponsored a study and supplied the
necessary equipment to determine horizon defi-
nition as applied to the design of navigation
and guidance systems. A few mosaic prints
were derived from a series of exposures taken
of the horizon. The MIT photographic study
is discussed, and a sample photograph is shown
in the Pilot Performance paper (paper 6) .

A meteorological experiment involving a
series of special photographs for the U.S.
Weather Bureau was not accomplished because
of the lack of time.

Astronaut Carpenter exposed an extensive
series of general interest color photographs of
subjects ranging from terrestrial features and
cloud formations to the launch-vehicle tankage
and the tethered balloon. Some of these photo-
graphs are displayed in the Pilot's Flight Re-
port (paper 7).

Launch Vehicle Performance

The launch vehicle used to accelerate Astro-
naut Carpenter and his Aurora 7 spacecraft
into orbit was an Atlas D missile modified for
the Mercury mission. The MA-7 launch vehi-
cle was essentially the same as that used for the
MA-6 mission and is described in paper 4 of
reference 1.

The differences between the Atlas 107-D
(MA-7) and the Atlas 109-D used for MA-6
involved retention of the insulation bulkhead
and reduction of the staging time from 131.3 to
130.1 seconds after lift-off. The performance
of the launch vehicle was exceptionally good,
with the countdown, launch, and insertion con-
forming very closely to planned conditions. At
sustainer engine cutoff (SECO), all spacecraft
and launch-vehicle systems were go, and only
one anomaly occurred during launch which re-
quires mention.

Although the abort sensing and implementa-
tion system (ASIS) performed satisfactorily
during the flight, hydraulic switch no. 2 for the
sustainer engine actuated to the abort position
at 4:25 minutes after lift-off. This switch and
the pressure transducer H52P for the sustainer
hydraulic accumulator are connected to a com-
mon pressure-sensing line as shown in figure
1-14. For an unknown reason this transducer
was apparently faulty and showed a gradual

Control system return

(switch j

\n ol7

_|low press.;
| mom fold 1

! i

1 Reservoir!

High press.;

mcnifold |

L_ . J

Figure 1-14. — Launch- vehicle hydraulic diagram.

decrease in pressure from 2,940 psia to 0 between
190 and 312 seconds after lift-off. Another
transducer located in the sustainer control cir-
cuit indicated that pressure had remained at
proper levels throughout powered flight;
therefore, this pressure switch did not. actuate
until the normal time after SECO. Since both
of these switches must be activated to initiate
an abort, the signal which would have unnec-
essarily terminated the flight was not generated.

Reference

1. Anon: Results of the First U.S. Manned Orbital Space Flight, Feb. 20, 1962. XASA Manned Spacecraft
Center.

2. MERCURY NETWORK PERFORMANCE

By James J. Donegan, Manned Space Flight Support Division, NASA Goddard Space Flight Center ; and
James C. Jackson, Manned Space Flight Support Division, NASA Goddard Space Flight Center

Summary

The Mercury Network performed very well
in support of the Mercury Atlas-7 mission. All
systems required to support the mission were
operational at the time of launch and in some
instances utilized backup equipment in the place
of primary equipment. No problems were en-
countered with computing and data flow. The
computers at the Goddard Space Flight Center
accurately predicted the 250 nautical mile over-
shoot immediately after the FPS-1G tracking
data from Point Arguello, California, were re-
ceived. Radar tracking was generally horizon
to horizon, and the resulting data provided to
the Goddard computers resulted in good orbit
determination during the mission.

The ground communications network per-
formance was generally better than that of the
MA-G mission. The ground-to-spaeecraft com-
munications were slightly inferior to MA-G per-
formance, particularly when patched onto the
conference network to allow monitoring by
other stations. Telemetry reception, as in the
MA-6 flight, was good.

Introduction

The purpose of this paper is twofold. The
first is to present a description and the per-
formance during the MA-7 mission of the Mer-

cury Network and its associated equipment.
The second is to describe briefly the Mercury
real-time computing system of the network and
to give a brief account of its performance dur-
ing the MA-7 mission.

Mercury Network

The Mercury Network configuration for the
MA-7 flight shown in figure 2-1, was the same
as that for MA-6, with but minor exceptions.
For MA-7 there was no Mid-Atlantic Ship, the
Indian Ocean Ship was repositioned in the
Mozambique Channel as shown in figure 2-1.

The Mercury Network consists of 15 Mercury
sites supplemented by several Atlantic Missile
Range (AMR) stations, and the Goddard
Space Flight Center communications and com-
puting facility. The major functions of this
Network during the MA-7 mission were to :

(1) Provide ground radar tracking of the
spacecraft and data transmission to the God-
dard computers.

(2) Provide launch and orbital computing
during the flight with real-time display data
transmitted to the Mercury Control Center.

(3) Provide real-time telemetry display data
at the sites and summary messages to Mercury
Control Center (MCC) for flight control pur-
poses.

(4) Provide command capability from vari-
ous stations for astronaut backup of critical
spacecraft control functions.

(5) Provide ground-to-spacecraft voice com-
munications and remote station-to-MCC voice
and teletype communications.

The major equipment subsystems located at
each site are shown in table 2-1. Generally,
the overall performance of all major equipment
of the Mercury Network during the MA-7 mis-
sion was excellent. A brief description of the
performance of each specific network subsystem
and an introduction to the equipment is pre-
sented in the following sections.

Radar Tracking and Acquisition

Two principal types of precision tracking
radars are used in the Mercury ground range
to track the spacecraft : the AN/FPS-16 and
Verlort radars. The AN/FPS-16, shown in
figure 2-2, is a precision C-band tracking radar
with a 12-foot dish. It operates on a frequency
of 5,500 to 5,900 me and has a beam width of
approximately 1.2°. It is the most accurate
of our tracking devices. The S-band, or Ver-
lort, radar, shown in figure 2-3, is a very long-
range tracking radar with a 10-foot dish. It
operates on a frequency of 2,800 to 3,000 me and
has a beam width of approximately 2.5°. The
redundancy provided by both of these radar
systems supplies the computers with sufficient
data to determine the orbit, should one of the
spacecraft beacons fail. The active acquisition
aid has a quad-helix antenna, which is shown
in figure 2-4. It has a broad beam width of 20° ,
operates on telemetry frequencies (215 to 265

Figure 2-2. — AN/FPS-16 radar installation at
Bermuda.

Figure 2-3. — Verlort radar installation at Guaymas,
Mexico.

me), and normally acquires the target first.
The acquisition aid console and equipment racks
are shown in figure 2-5.

This acquisition capability is most critical
at sites with the FPS-16 radar since this radar
is a narrow -beam device requiring precise point-
ing information to locate the target. Once the
yadar has acquired the spacecraft, the radar

Figure 2-4. — Acquisition aid quad-helix antenna at
Bermuda.

Table 2 - 1 . — Grov/nd Cornmunhations

Ground Com-

Com-

Telem-

Air-

Acqui-

munications

Station

Orbital-pass

coverage

mand

etry re-
ception

ground

voice

FPS-16

Verlort

sition

Computer

Timing

control

radar

aid

Voice

Telem-

etry

Mercury Control Center (MCC)

B/GE \

1P7090 J

1, 2, and 3. ,
1, 2, and 3_.
1, 2, and 3..

Grand Turk (GTI)*

Bermuda (BDA)

IBM-709

1, 2, and 3__

Canton Island (CTN)

2 and 3

1, 2, and 3__

White Sands, N M (WHS) b

Eglin Florida (EGL)*>

MPQ-

Goddard Space Flight Center
(GSFC).

IBM-7090

Comm.

Center, _

» No monitoring facilities; downrange antennas for MCC.
h Radar tracking station only.

Figure 2-.". — Acquisition aid console and equipment
racks.

system begins automatic tracking and does not
require additional acquisition assistance unless
the (racking is interrupted.

The aequisit ion system performance was very
good; the only difficulty encountered was the
failure of the elevation drive motor at the Zan-
zibar station. This failure did not influence
the reception of data since the operator was aide
to operate successfully the antenna elevation in
the manual mode. The coverage periods of the
acquisition system for the Xetwork are shown
in figure 2-0.

A comparison of the radar coverage for
MA-fi and MA-7, for both C- and S-band sys-
tems, is shown in figure 2-7. From an examina-
tion of this figure, it can be ascertained that
the acquisition system received signals beyond
the meaningful limits of horizon-to-horizon
track. The standard errors in range, azimuth,
and elevation as a result of noise in the radar
data collected by the Goddard computers and
the quantity of data received are given in table

2-1 T for the IMA -7 flight. These data reveal
that the radar tracking was comparable with
the horizon-to-horizon coverage obtained dur-
ing MA-fi.

Tracking was consistent from horizon to hori-
zon. The spacecraft beacons functioned very
well during the launch phase and satisfactorily
throughout the flight. Some amplitude modu-
lation and slight beacon countdown were noted
at times. However, these conditions caused no
noticeable deterioration of data presented to the
computers. Signal strengths received by the
radars were noticeably weaker than in the
MA <> flight. The radar data transmission (via
automatic teletype) was excellent with only
minor errors in transmission of several lines of
data from Muehea, Australia, during the first
orbital pass. There was a total of 07,354 char-
acters transmitted by the network radars with
no error.

Computing

Ily way of introduction to the Mercury real-
time computing system, a brief description is
given. Data from the worldwide Mercury
Tracking Xetwork are transmitted to the God-
dard Communications Center via the data cir-
cuits shown in figure 2-8. From the Communi-
cations Center, the data are transmitted to the
Goddard Computing Center, shown in figure
2-9, which is located in an adjacent room.
Here, real-time equipment places the radar data
from each tracking station automatically in
the core storage of the computers. Two IBM
7090 computers operating independently but in
a parallel fashion process the data. Should
a computer malfunction during the mission, the
other computer may be switched on-line to sup-

Figure 2-7. — Comparison of radar coverage for MA-6 and MA-7.

Vancouver

port the mission while the malfunctioning com-
puter is taken off-line and repaired.

The Mercury computing program is a real-
time automatic computing program designed to
provide trajectory information necessary to the
flight control of the Mercury mission. The
heart of the real-time computing system is
the monitor system which is shown schemati-
cally in figure 2-10. This monitor control sys-
tem directs the sequence of computer operations
in real time. Simply stated, the monitor system
accepts data from the remote sites, places the
data in the correct block of computer memory,
calls on the correct processor (whether it be

Figure 2-9. — Goddard Computing Center.

Table 2-II. — Standard, Deviations of MA-7 Low-Speed Radar Data

Station

Radar

Total

points

Standard deviations

Range, yards

Azimuth, mils

Elevation, mils

First orbital pass

FPS-16

31. 0

0. 11

0. 54

Verlort .

62. 8

1. 28

2. 15

Verlort

18. 5

1. 18

. 67

Muchea --

Verlort ..

17. 8

1. 05

1. 15

FPS-16.

4. 5

. 17

. 14

Verlort

11. 0

1. 44

1. 58

FPS-16

4. 9

. 23

. 41

Texas

Verlort

31. 9

2. 55

2. 20

FPS-16

10. 0

. 35

. 24

Verlort

40. 1

1. 67

1. 78

Cape Canaveral -

FPS-16

7. 6

. 12

. 56

Second orbital

pass

FPS-16

10. 1

0. 16

0. 57

Verlort

62. 2

1. 65

2. 71

Verlort .

12. 0

2. 31

1. 80

Verlort..

22. 9

1. 28

1. 34

FPS-16

2. 5

. 10

. 22

Hawaii

FPS-16

5. 4

. 24

. 21

Verlort.. _

16. 7

1. 33

1. 23

FPS-16

10. 1

. 30

. 40

Verlort..

12. 0

1. 42

1. 56

FPS-16

17. 7

. 14

. 39

Verlort..

76. 7

2 . 50

2. 44

FPS-16

7. 0

. 54

. 29

FPS-16

89. 1

I. 82

2. 53

Cape Canaveral

FPS-16

6. 6

. 16

. 73

Third orbital pass

FPS-16

8. 2

0. 30

0. 50 -

Verlort .

34. 7

1. 84

2. 09

Muchea__ __ _

Verlort .

8. 8

. 74

. 60

Hawaii

FPS-16

11. 5

. 19

. 34

Hawaii. _ -

Verlort . . .

21. 8

1. 83

1. 71

Reentry

California-

FPS-16

11. 6

0. 84

0. 66

Verlort

20. 5

1. 93

1. 64

FPS-16

21. 7

. 23

1. 14

Verlort

91. 6

2. 37

2. 19

Verlort

39. 4

1. 17

. 32

Cape Canaveral

FPS-16.

10. 0

. 14

. 43

San Salvador.. . .

FPS-16

30. 8

. 21

. 36

Reentry

computations

Data <
•ntry

Monitor 1

control system 1
{Input,
priority, etc.)

Which _
routine?

Orbit

computations

Monitor

control system -
(output)

. Data
exit

Launch

programs

Figure 2-10. — Real-time monitor control system.

launch, orbit, or reentry) to perform the proper
computation on the data, then provides the re-
quired output quantities to be transmitted to
the proper destination at the correct time.

During the MA-7 mission the computing
system at Goddard performed well. The
equipment, the launch subsystem, and the high-
speed line functioned properly during the en-
tire mission. Especially gratifying was the
performance of the Bermuda high-speed data
system and computations which were imple-
mented after the MA-G mission. The new
dual-compilation system also worked well.

Launch . — All the computing and data trans-
mission equipment was operational during the
entire countdown. High-speed input data were
continuous during the powered phase of the
flight from each of the three data sources, the
AMR range safety computer, the launch-vehi-
cle guidance computer, and the Bermuda range

Taiu.e 2-TTI. — Launch Phase Discrete and
Telemetry Events

Event

Time "tag” of arrival at
GSFC, sec since lift-off
unless indicated

General Electric/
Burroughs line

Nominal

7:45:16 e.s.t

BECO

128.964. _ ... _.

130. 1
152. 2
142. 0
304. 7
306. 3

Tower release. .
Tower separation __
SECO . _

153.464. . __

153.464. _

310.464

Spacecraft separa-
tion.

313.964. _

317.964.

318.464. .. . ...

318.964

Orbit-phase switch.

350.964 . -.

station. See figure 2-11. The data received
from Atlantic Missile Range sources during the
launch were excellent. At lift-off FPS-16
data processed through the AMR IP 7090 com-
puter were used as the data source for the God-
dard computers for approximately the first 35
seconds of launch. Mark II Azusa data pro-
cessed by the AMR IP 7090 computer were
used for the next 37 seconds as source date for
the Goddard computers. The launch-vehicle
guidance complex acquired the vehicle in both
rate and track at 00:01: 02 g.e.t. and was used
throughout the powered-flight phase, and during
the “go-no-go” computation as the selected data
source by the Goddard computers. Minor de-
viations in flight-path angle, for example, 1.2°
at booster-engine cutoff, and altitude during
powered flight were corrected by steering prior
to insertion by the Atlas guidance system.
Time of the telemetry discretes observed by
Goddard during launch are shown in table 2-
III. Insertion conditions computed on the
basis of the three independent launch sources
of data were in close agreement. The data
from all sources during the launch were excel-
lent. From a trajectory point of Hew it was a
nearly perfect launch.

Orbital phase . — As a result of the extremely
good insertion conditions provided by the Atlas
launch vehicle, the orbit phase was nearly nom-
inal. The orbit was determined accurately and
verified early in the first pass. The orbital
computation equipment functioned normally
and automatically during the mission. A basic
parameter which is usually indicative of the
performance of the tracking-computing net-
work is the computed time for retrorocket igni-
tion for a landing in the normal mission recov-
ery area. This parameter varied a maximum
of 2 seconds from launch throughout the mis-
sion after the Bermuda correction. As stated
previously, the tracking data were plentiful
and accurate during orbit.

Reentry phase . — A retrofire time of 4 hr
32 min 58 sec g.e.t. was recommended by the
Goddard computers based on a nominal land-
ing point of 68° W. longitude. The retrofire
time actually used was 4 hr 33 min 06 sec g.e.t.
based on a more realistic reentry weight because
of the actual fuel usage. The retrorocket s were
fired at approximately 04 :33 :09 g.e.t. Point
Arguello, Calif., tracked the spacecraft during

Bermuda

Figure 2-11. — Computer data sources during- launch.

and after retrofire. Based on the Point Ar-
guello FPS-16 tracking information, the God-
dard computers immediately predicted an over-
shoot of 246 nautical miles. The overshoot
point was confirmed by the data from the 'White
Sands and Texas stations and all subsequent
tracking data. The position of the spacecraft
was continuously and accurately displayed on
the wall map of the Mercury Control Center in
real time to an altitude of 60,000 feet.

From an analysis of the data, it appears the
tracking and computing systems performed
their primary tasks normally and without ex-
ception. No computer or equipment problems
were encountered during the mission.

Telemetry and Timing

The telemetry system provides reception of
the aeromedical data for display of astronaut
heartbeat rate, respiration, ECG, blood pres-

Figure 2-13. — Typical telemetry receiving equipment

Figure 2-13. — Continued.

sure, and body temperature. It also provides
the reception and display of data indicative of
spacecraft performance for use by the Flight
Control team at each tracking station.

A typical telemetry antenna installation is
shown in figure 2-12, and the associated elec-
tronic receiving and decommutation equipment
is shown in figure 2-13. A typical arrange-
ment of display consoles for Flight Control at
remote sites is shown in figure 2-14. Although
not all spacecraft systems quantities are dis-
played at the Flight Control consoles, all data
received are recorded either on magnetic tape
or on direct-writing oscillograph recorders.
The timing system provides time marks on all
records for later verification and also provides

time “tags” with the radar data transmitted to
the computers.

In general, all tracking stations received
telemetry signals from horizon to horizon. Be-
cause the telemetry transmission frequency is
severely attenuated by the reentry ionization
sheath, a blackout of ground reception results.
This effect was recorded for MA-7 as com-
mencing at a ground-elapsed time of 4 hours,
43 minutes and 58 seconds. Signal contact was
regained at 4 hr 48 min and 47 sec for approxi-
mately 12 seconds at Grand Turk Island.
Final loss of telemetry during the landing
phase was the result of extreme range and low
elevation angle.

A comparison of telemetry reception cover-
age for each site for MA-6 and MA-7 is given
in figure 2-15 and the received signal strengths
are given in table 2-IV. The reception periods
for each station, identified in figure 2-15, are
almost identical to the results shown for the
MA-6 mission.

Command

The dual ground command system was in-
stalled at specific command sites shown in table
2-1. This system provides ground command
backup to critical spacecraft functions such as
abort and retrofire. Out of a total of 16 com-
mand functions transmitted to the spacecraft,
15 were effectively received. The one exception
was a calibration command transmitted from
Muchea, Australia, which was attempted when
the spacecraft had passed beyond the optimum
transmitting time.

As backup means of voice communications
from the ground to spacecraft, the ground com-

Figure 2-14. — Typical display consoles at remote
stations.

Table 2-1 V. — Telemetry Receiver Signal Strengths

Station

Estimated mean, microvolts

Low (receiver
1, model 1415)

Low (receiver
2, model 1434)

High (receiver
1, model 1415)

High (receiver
2, model 1434)

First orbital pass

Bermuda . .-

100

Canaries

275

Kano - - .

150

Zanzibar -

Indian Ocean Ship -- - —

Muchea - -

195

175

195

180

Woomera —

300

250'

Canton

100

120

100

150

Hawaii,- -

Not in range

Guaymas

California

Texas —

300

195

300

Second orbital pass

Bermuda _ - - -

150

100

110

Canaries —

110

170

Kano

Zanzibar

Indian Ocean Ship- --

200

Muchea- ---

195

185

180

165

Woomera - - - -- —

300

190

160

120

Canton

Hawaii -

100

Guaymas --- .

California --

Not recorded

Texas. -

Third orbital pass

Bermuda —

Canaries —

Too low to estimate

Kano _ _ .-

Not in range

Zanzibar . .. -

Not in range

Indian Ocean Ship. . ..

120

200

Muchea. .

135

100

150

Woomera --

Canton - - -

Not in range

Hawaii.. --

200

100

200

Guaymas - .....

California ..

Not recorded

100

Texas

200

300

325

Figure 2-15. — Comparison of telemetry coverage times for MA-6 and MA-7.

mancl system employs a voice modulator which
may be utilized in the event both the HF and
UHF voice systems are inoperative. This back-
up technique was tested successfully during the
first orbital pass over Muchea, Australia.

Standby systems were called upon to main-
tain command coverage at two stations. Minor
trouble was experienced at the California sta-
tion during the first orbital pass when a fuse
opened in the primary system, and the trans-
mitter in the primary command system at Guay-
mas failed during the third pass. In both cases
the standby system functioned satisfactorily.
The command system operated normally in
spite of several minor malfunctions and had
no effect on the mission.

Communications

The communications facilities for the Mer-
cury Network consist of :

(1) Teletype between all remote stations and
the Mercury Control Center through the
Goddard Space Flight Center.

(2) Direct-line telephone communications
between selected stations and Mercury
Control Center.

(3) HF and UHF communications from all
stations except Eglin, Fla., and White
Sands, N. Mex., to the spacecraft.

The teletype circuits are utilized for flight-
control message traffic and radar data from
radar tracking stations to the Goddard com-
puters. This sharing of circuits restricts all
but priority messages from a station during
the radar tracking period.

The voice communications circuits between
the Mercury Control Center and selected re-
mote stations provide a direct and rapid means
of information transfer between the Flight
Control personnel at these locations.

The HF and UHF receiving and transmit-
ting equipment permits direct and successive
voice contact during the flight between the as-
tronaut and the Flight Control team over each
station. A patching arrangement permits all
stations which have ground voice communica-
tions to the Control Center to monitor all
ground-to-spacecraft communications during
the flight.

Teletype and ground voice . — All regular,
part-time, and alternate circuits were active and
operative on launch day. The propagation pre-
diction was for good conditions for those tele-
type circuits which utilize radio links to reach
certain stations. The teletype network per-
formed well with only three difficulties occur-
ring:

(1) The teletype circuit between Goddard
and Guaymas was open for a 7-minute
period beginning at 00 :36 g.e.t. The
spacecraft was not over Guaymas at the
time, and retransmission assured message
continuity.

(2) The Australian cable gave trouble for a
short period at 01 : 00 g.e.t., which re-
sulted in loss of four lines of Woomera
radar data.

(3) Teletype traffic to the Indian Ocean Ship
was interrupted for about 6 minutes due
to propagation; however, the interrup-

■-

-d

■

>/U

period J

Kano

Zonzibor

...

—

location o

hip

banged

A-

—

■

“

1—

fed

Texas

—

Canaveroi

—

■

Grand Turk Is.

■

Bermuda

■

Atlantic Ocean 9tip

Not

jsed on N

ilA-7

Canaries

■

Indian Ocean Ship

occ

iti

hif

hanged

MA-

-C

Canton

1—1

”

]

Californio

■

Texas

t i

Eqlin

...

(-Blackout)

period

0 00

030

100

130 200

Ground elapsed time, hr

200

min

300

330

400

4 30

UHF

Figure 2-16. — HF and UHF coverage for MA-6 and MA-7.

tion occurred at a (ime when tlie space-
craft was approaching the west coast of
the United States and did not interrupt
critical traffic.

II F and TJIIF voice . — The quality of the
ground-to-spacecraft communications was ac-
ceptable throughout the mission; however, it
was not as good as that for the MA-6 mission.
A study of the character of the average signal
strength at the ground systems reveals that the
majority of the stations reported a lower signal
level for MA-7 than was experienced during

MA-6. In some of these cases the signal level
was 2 to 5 times greater for the earlier mission.

It was noted that when the ground-to-space-
craft circuit was patched to the between-sta-
tions voice conference, the quality was not as
good as the MA-6 mission. This effect is being
investigated by studying the recordings made at
various locations on the circuits. The general-
ly weaker signal strength may account for part
of this problem. Figure 2-16 shows the HF
and UIIF coverage for both the MA-6 and
MA-7 missions.

3. MISSION OPERATIONS

By John D. Hodge, Asst. Chief for Flight Control, Flight Operations Division, NASA Manned Spacecraft
Center ; Eugene F. Rranz, Flight Operations Division, NASA Manned Spacecraft Center; and Wil-
liam C. Hayes, Flight Operations Division, NASA Manned Spacecraft Center

Summary

A discussion of the detailed operational sup-
port provided during the MA-7 mission, in-
cluding prelaunch, launch, flight, and recovery
phases, is presented. Since the launch vehicle
countdown and prelaunch phase was nearly
identical to that for MA-6, this activity is given
only minor emphasis. The launch phase pro-
ceeded almost perfectly, with only a last-minute
hold for weather. Powered flight was normal,
and the Mercury spacecraft was inserted into
a nominal orbit with exceptional precision.
The flight was satisfactorily monitored by the
ground stations of the Mercury Network, and
their activities are presented chronologically.
No major flight discrepancies were evident dur-
ing the orbital phase until just prior to retro-
fire, when it was discovered that the automatic
control system was not operating properly.
The astronaut was instructed by ground per-
sonnel to effect a manual retrofire maneuver.
Radar tracking data subsequent to this maneu-
ver indicated that the spacecraft would land
250 nautical miles downrange of the planned
landing point. Following contingency recov-
ery procedures, the astronaut was recovered by
helicopter some 3 hours after landing, and the
spacecraft was retrieved by a recovery destroyer
approximately 3 hours later.

Introduction

In the present paper, the flight control and
recovery operations for the MA-7 mission will
be discussed in detail. Since the launch sup-
port procedure was discussed in the MA-6
flight report (ref. 1), it will be only discussed
briefly. Some small changes from the MA-6
operational support were made, most of which
were associated with the development of ap-

propriate support procedures for the future
missions of longer duration. Network support
is discussed in detail in paper 2. Based on
previous experience, it was found that the total
recovery support used for MA-6 could be
slightly reduced for MA-7. The flight plan
was basically the same as that for the MA-6
mission with two significant differences: the
astronaut was given a greater amount of man-
ual-control tasks to perform ; and a large num-
ber of experiments were to be accomplished.

Prelaunch Activities

During the prelaunch period, four flight con-
troller network exercises were performed. A
new network countdown was used, and a high
degree of confidence was established in the
countdown format during this time. These
drills were very similar in content to those per-
formed for the MA-6 mission. Flight Con-
trollers quickly obtained a high degree of con-
fidence in various site and network procedures
and reached a high level of performance very
early in the schedule. They maintained this
performance level throughout all the network
simulations and during the actual MA-7 flight.
The countdown for launching the Mercury-
Atlas vehicle is conducted in two parts. The
first part is conducted on the day before the
launch and lasts approximately 41/2 hours. This
part of the countdown was conducted with no
major problems or delays. The second part of
the countdown was probably as close to per-
fect for the launch vehicle, spacecraft, and net-
work as could ever be expected. There were
some minor problems throughout the network ;
however, none of these resulted in the necessity
for a hold, and the coordination between the
various agencies involved was excellent. A hold
of about 45 minutes occurred at T-ll minutes

654533 0—62 3

in anticipation of better camera coverage and
to allow aircraft to check the atmospheric re-
fraction index in the vicinity of Cape Canaveral
for the launch-vehicle guidance equipment.

Powered Flight Phase

The launch occurred at 07 : 45 : 16 a.m. e.s.t. on
May '24, 1962. Sustainer engine cutoff occurred
at 5 minutes 10 seconds ground elapsed time
(g.e.t.). The “go” capability as indicated by
the Goddard Space Flight Center computers
was obtained and transmitted to the astronaut
at 5 minutes 32 seconds. The powered portion
of flight was completely normal, and the astro-
naut was able to make all of the planned com-
munications and observations throughout this
period. The Mercury Control Center go-no-go
decision at cutoff was made rapidly, and there
was no doubt that conditions very close to no-
minal had been achieved. Table 3-1 presents

the actual cutoff conditions that were obtained.
A comparison of the planned and actual times
at which the major events occurred are given
in table 3— II.

Table 3-1. — Actual Flight Conditions

Cutoff conditions :

Altitude, ft 527, 859

Velocity, ft/sec 25, 717

Flight-path angle, deg —0. 0004

Orbit parameters :

Perigee altitude, nautical miles 86. 87

Apogee altitude, nautical miles 144. 96

Period, min : sec 88 : 32

Inclination angle, deg 32. 55

Maximum conditions :

Exit acceleration, g units 7. 8

Exit dynamic pressure,* lb/sq ft 967

Entry acceleration, g units 7. 5

Entry dynamic pressure, lb/sq ft 429

* Based on the atmosphere at Cape Canaveral.

Table 3-IT. — Sequence of Events During MA-7 Flight

Event

Preflight predicted
time, hr:min:sec

Actual time,
hr:min:sec

Booster-engine cutoff (BECO)

Tower release..

00:02:10.1

00:02:32.2

00:02:08.6

00:02:32.2

00:05:09.9

00:05:05.3

00:05:10.2

Spacecraft separation

00:05:06.3
04:32 :25.6

00:05:12.2

04:32:36.5

04:32:55.6

04:33:10.3

04:33:00.6

04 :33:15.3

04:33:05.6

04 :33:20.5

04:33:55.6

04:34:10.8

04:43:55.6

04:44:44

04:50:00.6

04:50:54

04:50:37.6

04 :5 1:48.2

04:55:22.6

04:55:57

04:55 :22,6

04:56:04.8

Orbital Flight Phase

After separation of the spacecraft from the
launch vehicle, the astronaut was given all per-
tinent data involved with orbit parameters and
the necessary retrofire times were transmitted.
A remoting facility for transmitting air-to-
ground voice for the Mercury Control Center
through the Bermuda site transmitters was im-
plemented for the MA-7 mission. This facility
enabled the Mercury Control Center Capsule

Communicator (Cap Com) to transmit space-
craft. systems data and orbital information to
the astronaut in real time; therefore, much
of the requirement for relaying information be-
tween the Canaveral and Bermuda Flight Con-
trollers was eliminated. From the summary
messages received from the African sites, it
became readily apparent that the suit cooling
system was not correctly adjusted and that the
astronaut was uncomfortable. However, the

suit temperature began to decrease when the
astronaut increased the water flow in the suit
cooling circuit. By the end of the first orbit
it had reduced to a satisfactory value. Other
than the slight discomfort due to the high suit
temperature, the astronaut was obviously in
good condition and performing satisfactorily
throughout the first orbit. The Canary Island
site transmitted radar data to the Goddard com-
puters, and these data confirmed the orbital in-
sertion parameters and an extremely good or-
ital definition was obtained. Over the Woo-
mera station, the astronaut reported that he
took four swallows of water and that his bite-
sized food tablets had crumbled in the container
and some particles of food were floating free in
the cabin. He was able to eat some of the
crumbled food.

Toward the end of this pass, a slight in-
crease in body temperature was noted. The
Canton Island site then reported a body tem-
perature of 102°. However, the Mercury Con-
trol Center surgeon felt such a rapid increase
was not probable and that the transducer had
either failed or had been affected by the tele-
metry calibrate command transmitted from the
Muchea site. The only other problem was the
large amount of automatic control system fuel
being used by the astronaut during the first
orbit. He was cautioned against further ex-
cessive usage of this fuel during the orbital
pass over the United States. The air-to-
ground transmissions relayed via the Goddard
voice loop during the first orbit were of good
quality and provided the Mercury Control Cen-
ter with information available from the air-
ground voice communications of the astronaut.
The network air-ground voice quality, although
not as good as the previous MA-6 mission, con-
tinued to be usable throughout the remaining
orbits and provided one of the best tools for
maintaining surveillance of the flight. The
spacecraft clock performed satisfactorily
throughout the entire mission. The initial
clock error of —1 second remained constant
throughout the mission and was compensated
for in the retrosequence settings that were
transmitted to the astronaut. During the first
orbit, the network radar systems were able to
obtain excellent tracking data and these data,
together with the data obtained at cutoff, pro-
vided very adequate information on the space-

craft position and orbit. As an example, the
retrosequence time computed at insertion was
changed only 11 seconds by the Bermuda data,
and thereafter, the time varied within only ±1
second throng] unit the mission. The balloon
was deployed during contact with Cape Cana-
veral 1 hour 38 minutes after lift-off. During
the first portion of the second orbit, the suit
temperature indicated a rise from 70° at Cape
Canaveral to approximately 90° during con-
tact with the Indian Ocean Ship, but again
showed a decrease in trend before acquisition
by the Muchea and Woomera stations. It was
obvious throughout the flight that the pilot was
having difficulty in achieving the proper water-
flow setting for the suit cooling system. There
is about a 30-minute lag in the cooling system
in response to a change in the valve setting and
as a result it was difficult to determine an ade-
quate setting. However, when the loss of sig-
nal (LOS) occurred at Woomera, the suit tem-
perature had decreased to approximately 82°
and during the remaining one and one-half orb-
its the suit temperature indicated a steady de-
crease to a value between 64° and 67°. Cabin-
air-temperature readings were slightly higher
with the MA-6 flight. A maximum tempera-
ture of 108° was monitored by the Hawaii sta-
tion near the end of the second orbit. This
temperature decreased and tended to stabilize at
about 100° during the remainder of this orbit
and the first portion of the third orbit.

Over the Woomera site, the astronaut re-
ported that he could temporarily change the
spacecraft attitude by moving his arms
and body. The mission continued normally
throughout the remainder of the second orbit.
The astronaut was behind the flight-plan sched-
ule by several items, and it was noted at Cali-
fornia acquisition that the astronaut had used
rather large amounts of manual fuel and was
down to approximately 42 percent as he began
the third orbit. The low automatic and manual
fuel quantities caused considerable concern on
the ground and resulted in a further request
to the astronaut to conserve his fuel in both
the automatic and manual systems. Site evalu-
ation of telemetry recordings during the first
and second orbits indicated considerable high
thruster activity. These indications generally
occurred while the astronaut was in the fly-by-
wire mode, and it appeared that he was employ-

ing high thrusters excessively during attitude
changes.

Throughout the flight, the astronaut made a
number of voice reports regarding visual ob-
servations and discussed various experiments
carried out in the flight. These reports are ex-
plained in more detail in paper 7.

The Mercury Control Center made a go deci-
sion for the beginning of the third orbit at 2
hours 55 minutes g.e.t. The astronaut was cau-
tioned to conserve his fuel and it was suggested
that he increase his water flow to the inverter
cold plates. The inverters had indicated an
increase in temperature similar to the previous
MA-6 flight. This caused no major concern;
however, the increased water flow reduced the
rate of this temperature increase to an accepta-
ble level. As a result of the request to conserve
fuel, the astronaut entered a period of drifting
flight at 3 hours 9 minutes g.e.t. while he was
in contact with Cape Canaveral. Over Mer-
cury Control Center during the third orbit, 45
percent of the fuel in the automatic system and
42 percent of the fuel in the manual system re-
mained. Over the Indian Ocean Ship, the as-
tronaut attempted to jettison the balloon man-
ually and reported that he was unable to ac-
complish this although the switch was cycled
several times.

During the third orbit, all systems appeared
to be normal. The clock was reset to 04 :32 :34,
the retrofire time for the end of the third orbit,
by the astronaut while in contact with the
Muchea station.

Reentry Phase

Upon contact with Hawaii at the end of the
third orbit, the astronaut was instructed to be-
gin his preret resequence check list and to revert
from his present manual control mode to the
automatic mode in preparation for retrose-
quenee. The retrosequence check list was started
but when the astronaut returned to automatic
control, he reported having trouble with this
system and, as a result, was unable to complete
the list. The capsule communicator at Hawaii
continued transmitting the remainder of the
preret resequence check list after loss of tele-
metry contact, and most of the transmission was
received by the astronaut. However, the
ground was unable to confirm that it had been

received because of the limited UHF range of
the spacecraft.

From both astronaut reports and telemetry
readouts during the periods in which the astro-
naut was using automatic control over remote
sites during the mission, it appeared that no
major difficulty was experienced while using
this system. The astronaut reported the auto-
matic stabilization and control system (ASCS)
to be performing satisfactorily on several oc-
casions. Although some differences between
horizon scanner outputs and spacecraft atti-
tudes had been noted, they were not considered
to be any reason for concern because of the con-
trol configuration at the time. Therefore, the
failure of the ASCS system to maintain proper
attitudes when engaged by the astronaut over
Hawaii was unexpected. When voice com-
munications were established with the Cali-
fornia station, the astronaut continued to have
problems on ASCS and, with advice from the
capsule communicator, elected to perform the
retrofire maneuver using manual control.
During this period the astronaut used a com-
bination of window reference, periscope, and
attitude displays.

The astronaut was directed to initiate retro-
fire manually and to bypass the attitude per-
mission circuit. The coundown was transmit-
ted from California, but it was apparent that
the retrofire had taken place several seconds
late. Initial reports from the astronaut indi-
cated that the attitudes had been held fairly
well during retrofire. The California station
reported that the velocity change indicated by
the integrating accelerometer was normal. The
radar data from California indicated an over-
shoot but the indication was suspected to be in
error because of previous reports. However,
as additional radar data became available from
other sites, it was obvious that the California
radar data were correct and that the landing
point would be approximately 250 nautical
miles beyond the planned position. Because of
the small amount of automatic fuel remaining
following retrofire and the complete depletion
of manual fuel, the astronaut was instructed
to use as little fuel as possible in returning the
spacecraft to reentry attitude and to conserve
the fuel for use during reentry. He was also
instructed to use the ASCS auxiliary damping

mode during the atmospheric reentry portion
of the flight.

Upon contact with Cape Canaveral just
previous to the loss of communications as a
result of ionization blackout, the astronaut was
queried as to his face-plate position. He indi-
cated that it was still open, and proceeded to
close it. The communication blackout occurred
about 40 seconds late, an occurrence which lent
further evidence to the longer reentry range
predicted by the radar. The astronaut was told
that his landing point would be long and would
occur at approximately 19°23' N., 63°51' W.
From this point no voice communications were
received from the astronaut; however a brief
period of telemetry data was obtained after
blackout. A number of communications were
attempted with the command voice system and
the normal UHF and HF voice system on the
chance that he might receive this information.
All operating C-band radars at Cape Canaveral
and San Salvador tracked the C-band beacon
until the spacecraft went below the local hori-
zon indicating that the spacecraft had re-
entered satisfactorily, and these radar data con-
tinued to predict approximately the same land-
ing point.

Recovery Operations

The operation of recovery forces for this mis-
sion was very similar to that for the MA-6 mis-
sion. Planned landing areas were established
in the Atlantic as shown in figure 3-1 to cover
aborts during powered flight and landing at the
end of each orbital pass. The disposition of

the recovery forces in the planned landing areas
is shown in table 3-III. Area II is the planned
landing area for the end of the third orbit, and
recovery in this area could be effected within
3 hours of landing. Special aircraft were pre-
deployed on a standby basis to locate the space-
craft and render pararescue assistance within
18 hours of landing at any point along the
ground track.

During the entire mission, all recovery forces
were informed of the flight progress by the Re-
covery Control Center. Shortly after the as-
tronaut began the third pass, an Air Rescue
Service SC-54 aircraft with a specially trained
pararescue team aboard was dispatched as a
precautionary measure from Roosevelt Roads,
Puerto Rico, and assigned a position at the
downrange end of landing area H. As soon as
the calculated landing position was established
about 250 nautical miles downrange of the cen-
ter of area H, all units in area H were instructed

Table 3-III . — Disposition of Recovery Forces in Planned Landing Areas

Area

aircraft

Search and
rescue
aircraft

Helicopters

Ships

Maximum
recovery
time, hr

A»._

3 to 6

1 ■>

* Launch site recovery forces consisted of 3 helicopters, several amphibious vehicles, and small boats,
b Launched as a precautionary measure when the astronaut began the third orbital pass.

Table 3-IV . — Chronological Summary of Post-Landing Events

e.s.t.,
hr: min

Elapsed time
from landing,
hr: min

Event

11:18 a.m.

As a precautionary measure, Air Rescue Service SC-54 was launched from
Roosevelt Roads, Puerto Rico, to take station on downrange end of Area
H. The SC-54 had specially trained pararescue team aboard.

12:22 p.m.

Retrorockets were ignited.

12:33 p.m.

Calculated landing position was reported as being 19°24' N. latitude, 63°53'
W. longitude. Air Rescue Service SA-16 (amphibian) was launched and
instructed to proceed to this point.

12:35 p.m.

All units in area H were proceeding to calculated landing position.

12:41 p.m.