Post Launch Memorandum Report for Mercury-Atlas No. 9(MA-9). Part 1 Mission Analysis

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(H1SA-TM-X-61634) E O S T L A O N C H M E M O R A N D U MH E F O B T F O B M E R C U R Y - A T I A S N O . 9 ( H A - 9 ) . P A R T1 : M I S S I O N A N A L Y S I S ( N A S A ) 3 5 0 p

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P O S T L A U N C H M E M O R A N D U M R E P O R T

F O R

M E R C U R Y - A T L A S N O . 9 (MA-9)(L0

PART I - MISSION ANALYSIS

|intervals; declassified

N A T I O N A L A E R O N A U T I C S A ND S P A C E A D M I N

M A N N E D S P A C E C R A F T C E N T E R

Cape Canaveral, Florida

June 2k, 1963

nent contains information affecting the national defense of the~MiHlMii. |-

I VJ • \J i *** » f

contents



POSTLAUNCH MEMORANDUM REPORT

FOR

MERCURY-ATLAS NO. 9 (MA-9)

PART I - MISSION ANALYSIS

Idited By: J. H. Boynton, Senior Editor

v R. G. Arbic

^C. A. Berry, M.D.

rfR. E. Day

P. C. Donnelly

,,W. R. Kelly

J. P. Mayer

Â. B. Shepard

R . E. Smylie

rt C M Scr t cS .H

O >s fl r-l 0

E H pq0o m

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

MANNED SPACECRAFT CENTER

Cape Canaveral, Florida

June 2k, 1963



TABLE OF CONTENTS

Section Page

1.0 MISSION SUMMARY 1-1

2.0 INTRODUCTION 2-1

3.0 SPACE VEHICLE DESCRIPTION 3-1

3.1 Spacecraft Description 3-1

3.2 Launch-Vehicle Description 3 - 1 5

I j - . O TRAJECTORY AND MISSION EVENTS U - 1

U.I Sequence of Flight Events U-l

U.2 Flight Trajectory U-3

5.0 SPACECRAFT PERFORMANCE 5-1

5.1 Spacecraft Control System 5-1

5.2 Life Support Systems 5-7

5-3 Communications Systems 5 - 2 75.U Mechanical and Pyrotechnic Systems 5 - 2 85-5 Electrical and Sequential Systems 5 - 3 05.6 Instrumentation System 5- 33

5.7 Heat Protection System 5 - 3^

5.8 Scientific Experiments 5 - 3 8

6.0 LAUNCH-VEHICLE PERFORMANCE 6-1

6.1 Airframe 6-16.2 Propulsion System 6-1

6.3 Propellant Tanking 6-2

6.U Propellant Utilization 6-2

6.5 Pneumatics ' . . . . 6-2

6.6 Electrical System 6-3

6.7 Flight Control System 6-3

6.8 Guidance 6-U

6.9 Abort Sensing and Implementation System 6-5

7.0 ASTRONAUT ACTIVITIES 7-1

7-1 Aeromedical Analysis 7-1

7.2 Pilot's Performance 7 - U-2

7.3 Pilot's Flight Report 7 - 7U

8.0 FLIGHT CONTROL AND NETWORK PERFORMANCE 8-1

8.1 Flight Control Summary 8-1

8.2 Mercury Network Performance 8 - 1 2



9.0 RECOVERY 9-1

9.1 Recovery Plans 9-1

9.2 Recovery Operations 9-2

9-3 Recovery Aids 9-3

0.0 APPENDIX A 10-1

10.1 Spacecraft History 10-1

10.2 Launch Procedure 10-6

10.3 Weather Conditions 10-7

10 A Flight Safety Review 10-9

10.5 Photographic Coverage 10-11

10.6 Postflight Inspection 10-11+

11.0 A P P E T O I X B - A C K W O W L E D G E M E E T 11-1



I

LIST O F T A B L E S

Table Page

3.1-1 SUMMARY OF SPACECRAFT SYSTEMS MODIFICATIONS 3-13

3.1-2 WEIGHT AND BALANCE DATA FOR SPACECRAFT 20 3 - l4

4.1-1 SEQUENCE OF EVENTS . . 4-2

4.2-1 COMPARISON OF PLANNED AND ACTUAL TRAJECTORY PARAMETERS . 4-5

5.1.5-1 MA-9FUEL USAGE 5-6

5.2.1.2-1 POSTFLIGHT ANALYSIS OF LITHIUM-HYDROXIDE CANISTER .... 5-25

5.2.4.1-1 BODY-MASS BALANCE SUMMARY 5-26

7.1.2.1-1 PILOT PREFLIGHT ACTIVITIES 7-24

7.1.2.1-2 PERTINENT EXCERPTS FROM CLINICAL EXAMINATIONS 7-25

7.1.2.1-3 COMPLETE BLOOD COUNTS 7-26

7.1.2.1-4 COMPARISON OF TYPICAL PREFLIGHT AND POSTFLIGHT URINE

VALUES 7-27

7.1.2.1-5 U R I N E A N A L Y S I S 7-28

7.1.2.1-6 L O W - R E S I D U E DIET 7-30

7.1.2.1-7 A E R O M E D I C A L C O U N T D O W N 7-31

7.1.2.2-1 D E T A I L E D P R E F L I G H T H E A R T - R A T E A N D RESPIRATION- R A T E

DATA 7-32

7.1.2.2-2 S U M M A R Y O F H E A R T - R A T E A N D R E S P I R A T I O N - R A T E DATA 7-3

7.1.2.2-3 D E T A I L E D P R E F L I G H T B L O O D - P R E S S U R E DATA 7 - 3 4

7.1.2.2-4 S U M M A R Y O F B L O O D P R E S S U R E DATA 7-35

7.1.3.1-1 S U M M A R Y OF C A L I B R A T E D W O R K 7-36

7.1.3.2-1 INFLIGHT SLEEP P E R I O D S 7-37

7.1.4.1-1 PILOT P O S T F L I G H T ACTIVITIES 7-38

7.1.4.2-1 R E C O R D OF'PILOT'S W E I G H T CHANGES 7-39

7.1.5.1-1 S U M M A R Y OFTILT S T U D I E S . 7 - 40



•GOIinDDMlAL •

Table Page

7.1.5.3-! BLOOD CHEMISTRIES 7 - 4l

7.2.1.1-1 PILOT TIME IN SPACECRAFT 20 DURING HANGAR AND LAUNCH

COMPLEX TESTS 7-59

7.2.1.2-1 PILOT TRAINING SUMMARY IN THE MERCURY PROCEDURES

TRAINER NUMBER 2 AT CAPE CANAVERAL 7 - 60

7.2.1.3-1 FLYING TIME FROM JANUARY 1 TO LAUNCH DATE 7 - 6l

7.2.1.5-1 PILOT PREFLIGHT ACTIVITIES FROM JANUARY 1, 1963 TO

LAUNCH DATE 7-62

7.2.2.1-1 SUMMARY OF MAJOR FLIGHT ACTIVITIES 7-65

7.2.2.2-1 PILOT'S EQUIPMENT LIST 7-71

7.2.3.2-1 SUMMARY OF ATTITUDE MANEUVERS 7-72

7.2.4.1-1 CONTROL MODE USAGE 7-73

8.2.1-1 ORBITAL INSERTION CONDITIONS AVAILABLE AT MCC 8-20

8.2.2-1 COMMAND HANDOVER SUMMARY 8-21

8.2.2-2 COMMAND FUNCTION SUMMARY 8-27

8.2.3-1 COMPUTER READOUT OF RADAR TRACKING DATA 8-28

8.2.3-2 RADAR TRACKING PERIODS 8-30

8.2.4-1 TELEMETRY COVERAGE 8-32

8.2.5-1 AIR-43ROUND COMMUNICATIONS COVERAGE 8 - 4l

9.1-1 RECOVERY SHIP AND AIRCRAFT DEPLOYMENT IN PLANNED

LANDING AREAS 9-5

10.5-1 AMR OPTICAL LAUNCH COVERAGE 10-13





G O l U r i D D M T l A L "

Figure Page

5.1-3-1 Postflight photograph of amplifier -calibrator

power socke't ... ................... 5-

5.1-3-2 Blue ribbon connector showing corrosion ......... 5-55

5.1.5-1 RCS auxiliary fuel tank ................. 5-56

5.1.5-2 Solenoid-inlet "B"-nut temperature -time history ..... 5-57

5.2.1.1-1 MA. -9 suit circuit condensate trap ............ 5-58

5.2.1.2-1 Cabin temperature evaluation ........... ^ . . . 5-59

5.2.2.1-1 MA-9helmet details ................... 5 - 60

5.2.3-1 Urine and condensate transfer system ........... 5 - 6l

5.2.3.1-1 Liquid transfer syringe ................ . 5-62

5.3-1 Onboard television system equipment ........... 5 - 63

5.3-5-1 TV picture of spacecraft interior ............ 5 - 6U

5-3.5-2 TV picture taken through the window ........... 5-64

5 - 4.3-1 Comparison of normal retropackage umbilical disconnect

squib with one missing main change ........... 5-65

5.6.1-1 l6-mm movie camera .......... -. ......... 5-66

5.6.2-1 MA-9 programer showing misalinement of faulty gear .... 5-67

5.7.1-1 Postflight photograph of MA-9ablation shield ...... 5-68

5.7-3-1 Postflight photograph of paint patches .......... 5-69

5.8.1-1 Camera for dim -light phenomena experiment . . ...... 5-70

5.8.1-2 Typical photograph taken for dim-light photography

experiment ....................... 5-71

5.8.2-1 Light assembly for ground light experiment ........ 5-72

5-8.3-1 Installation of flashing light and geiger counters on

retropackage ...................... 5-73



Figure Page

5.8A-1 Hasselblad 500-C'camera modified forMA-9mission ... 5 - 7^

5.8A-2 Infrared photograph taken for U.S.Weather Bureau

experiment over Florida peninsula 5-75

5.8.5-1 Quadrantal photograph taken for horizon-definitionexperiment » " V : • • 5-76

5.8.6-1 Equipment for flashing-light' experiment ..."...'.. 5-77

5.8.6-2 Topical history trace of flashing light output ; . . . . 5-78

5-8.6-3 Typical horizontal intensity distribution of theflashing light 5-79

5.8.6- Calculated sighting parameters for flashing-lightexperiment . ' ' . . ' . . . . 5 - 80

f t

5.8.8-1 Photograph'of Himalaya Mountains taken withHasselblad camera . ' •'. . . .5 -.8

6.8-1 Space-fixed velocity and flight-path angle in theregion of cut-off using launch-vehicle guidancedata

(a) Space-fixed velocity . . . . . 6-7

fb) Space-fixed flight-path angle . . 6-8

6.8-2 Space-fixed velocity and flight-path angle in t;heregion of cut-off using IP 709 data

(a) Space-fixed velocity . . ' , . . 6-9.

(b) Space-fixed flight-path angle . . . . . . . . . . '"6-10

6.-'8-5 Space-fixed flight-path angle versus space-fixedvelocity in the region of cut-off 6-11

7.1.2.2-1 Oral temperature probe 7-91

7.1.2.2-2 Installation of oral temperature probe in helmet . . . . 7-92

7.1.2.2-3 MA-9May l , 1963, 07: 2:00 e.s.t. Sample record

illustrating nodal beats occurring during canceledlaunch count down. Recorder speed 25 mm/sec ..'.'.. 7-93

C OPlPf f iENTlAL



C O N F I D E N T I A L

Figure Page

7.1.2.2-4 MA-9 16:11:30 - Sample of biosensor record at a rangestation illustrating one of the frequent occurrencesof sinus arrhythmia with wandering of the cardiac

pace maker. In this sample, the negative P wavesuggests inverse depolarization from the atrioventri-cular node. Similar changes were observed before

7.1.3.1-1

7.1.3.2-1

7.1.5.1-1

7.1.5.1-2

7.1.5.2-1

7.1.5.2-2

7.2.2.2*1

7.2.2.2-2

7.2.2.2-3

7.2.3.1-1

7.2.3.5-1

7.2.4.2-1

7.2.4.2-2

9.1-1

9.2-1

9.2-2

MA-9 12:29:52 - Sample of typical biosensor datareceived at a range station. Blood pressure,

Tilt studies - heart rate responses .....

Tilt studies - blood pressure responses for MA-9

Exercising device used for calibrated work

Calibrated work -MA-9. .

Special equipment storage kit

Planned landing areas

fa) Atlantic Ocean

(b) Pacific Ocean .

MA-9 spacecraft in auxiliary flotation collar with line

i

7

7

7

7

7

7

7

7

7

7

7

- 7

7

9

9

9

9

s .

- 95

-96

- 97

-98

- 99

- 100

- 101

- 101

- 102

- 103

- 104

- 105

- 106

- 6

- 7

- 8

- 9 "



iNTIAL

Figure • Bage

9.2-3 Side hatch being actuated on MA-9 spacecraft 9-10

9.2-4 Astronaut Cooper egressing from.MA-9 spacecraft .... 9 - H

9.3-1 Details of landing area 22-1 9-12

10.3-1 Wind direction and velocity at launch site 10-17

10.5-2-1 AMR engineering sequential tracking camera coverage . . 10-18

10.6-1 Postflight photograph of MA-9 spacecraft 10-19



N O T I C E

NO. 1: LIFT-OFF TIMS (2-INCH MOTION) FOR THE MA-9 FLIGHT WAS

8:04:13.106A.M.E.S.T. RANGE ZERO TIME WAS ESTABLISHED

AS 8:04:13 A.M.E.S.T. ALL TIMES REFERRED TO IN THIS

REPORT ARE IN ELAPSED TIME IN HR:MIN:SEC FROM RANGE ZERO

UNLESS OTHERWISE NOTED.

NO. 2: THE MA-9 POSTLAUNCH MEMORANDUM REPORT IS IN THREE PARTS,

UNDER SEPARATE COVERS, AS FOLLOWS:

PART I - MISSION ANALYSIS - CONTAINS AN OVERALL ANALYSIS

OF THE MISSION AND PRESENTS A MINIMUM OF DATA.

PART II - DATA - CONTAINS COMPLETE TIME HISTORIES OF

SPACECRAFT DATA, WITHOUT ANALYSIS.

PART III - MISSION TRANSCRIPTS - CONTAINS ESSENTIALLY

UNEDITED TRANSCRIPTS OF THE FLIGHT COMMUNICATIONS,,THE

PILOT'S POSTFLIGHT SELF-DEBRIEFING, AND THE FORMAL TECHNICAL

DEBRIEFING CONDUCTED ONBOARD THE RECOVERY AIRCRAFT CARRIER.

IDEMTVb.



AL Page 1-1

1.0 MISSION SUMMARY

The MA-9mission was successful in nearly every respect. The plannedlaunch time of 8:00 a.m. e.s.t. oh May l4, 19&3; was postponed for 1 day

because of intermittent digital data in both the azimuth and rangechannels of the C-band radar at Bermuda. Prior to postponement, the

countdown had proceeded as planned until T-60 minutes, when an unscheduledhold of 2 hours and 9 minutes became necessary because of a fuel-pumpfailure in the diesel engine on the gantry transfer table. After thishold, the countdown was continued until T-13 minutes when the flightwas postponed because of the radar problem. The launch operation on

May 15, 1963, was "the most efficient conducted to date. Four minutesof unplanned hold time were required to evaluate an external RF inter-ference problem at the guidance central rate station. Weather conditionsat the launch site and in the primary landing area were satisfactory.

Lift-off occurred at approximately 8:04 a.m. e.s.t. on May 15, 19 3

2 hours and 31 minutes after the astronaut entered the spacecraft.

Launch-vehicle performance was excellent, and the trajectoryparameters displayed at the Mercury Control Center indicated a "go"condition at insertion. A near-perfect orbit was attained, with deviations

from planned postposigrade values of space-fixed flight-path angle and

velocity of 0.0037° and -1.4 ft/sec, respectively. Both the perigeeand apogee of the initial orbit differed from the planned values of87 and 1^4 nautical miles by 0.2 nautical mile. The decay in perigee

and apogee after nearly 22 orbital passes was 1.6 and 7-1 nauticalmiles, respectively.

Spacecraft separation from the launch vehicle was satisfactory andthe planned manual turnaround was well executed by the pilot. Theperformance of the spacecraft systems was excellent for the first18 orbital passes with the exception that the automatic section of the

programer failed at 12:18:19- In addition, several minor problemswere encountered with the R and Z calibrations, the drinking-watervalve, and the condensate transfer system. Upon contacting Hawaii

on the 19th orbital pass, the pilot reported that the 0.05g warninglight had come on. Systems checks by the astronaut revealed that the

amplifier-calibrator was in the 0.05g configuration and that the ASCScould be used only during reentry. However, planned use of the ASCS

for reentry was abandoned at about 33:07:00 when neither the main nor

the standby 250 v-amp inverters would supply electrical power to theASCS bus. The pilot manually initiated the required retrofire andreentry events. He controlled the spacecraft attitudes during retro-fire by utilizing the manual proportional system. Because of the ASCSfailure, the pilot was also required to conduct the reentry maneuvermanually, and he elected to use both the manual proportional and fly-

by-wire modes during this phase.



P a g e i - 2

The pilot's performance throughout the mission was excellent, and

he adhered closely to the flight plan until the ASCS problems occurred.

The pilot had no difficulty in sleeping during the mission, although

he woke up several times during the planned rest period and found it

necessary to reestablish a comfortable suit temperature. He did not

eat and drink as much as was desirable, and he has since commented onthe difficulty of performing these functions with the devices that were

available to him. Six of the eight planned scientific experiments were

successfully conducted during the mission. The balloon drag and

visibility experiment was not accomplished because of failure of the

balloon to deploy, and the window attenuation experiment was not

accomplished because the pilot could not get the standard light source

out of the special equipment storage kit.

The pilot's control of the spacecraft during retrofire and reentry

was excellent and resulted in a landing only k.k nautical miles from

the prime recovery ship, the aircraft carrier U.S.S. Kearsarge. Visual

contact was made from the carrier and the. recovery helicopters reachedthe spacecraft and circled it during its descent. Swimmers were

deployed from the helicopters and they immediately attached a flotation

collar to the spacecraft. The pilot remained in the spacecraft until

it was hoisted aboard the carrier, the hatch had been blown, and the

doctors had given him a preliminary examination. The pilot egressed

from the spacecraft in good condition kO minutes after landing. A

postflight physical examination conducted onboard the recovery ship

revealed no evidence of significant degradation of pilot function directly

attributable to the space flight. The pilot demonstrated an orthostatic

rise in heart rate and fall in blood pressure which was more pronounced

than that detected after the MA-8 flight. Although this condition is

not an inflight hazard, the implications of this hemodynamic response

on return to Ig conditions will have to be given very serious consider-

ation for longer missions.

Support activities from all ground elements, including flight

control, Mercury Network, and recovery, were excellent and contributed

greatly to the successful accomplishment of the mission.

Postflight examination of the spacecraft and evaluation of the data

collected during the mission have revealed some anomalies, and detailed

systems tests have determined the most likely causes of the major

problems. Considerable information regarding man's capabilities to

perform his assigned tasks during extended periods of time in the space

environment has been obtained. Evaluation of the overall mission indicates

that a high degree of success was obtained and confirms the accomplishment

of all mission objectives.



2.0 INTRODUCTION

Page 2-1

The first manned 1-day mission (MOEM) as a part of the United

States' program of space exploration was successfully accomplished

on May 15 and May 1.6, 1963- This mission was the fourth mannedorbital flight in Project Mercury. It was also the ninth of a series

of flights utilizing production Mercury spacecraft and Atlas launch

vehicles and, therefore, was designated Mercury-Atlas Mission 9 (MA-9).

Astronaut L. Gordon Cooper, Jr., shown in figures 2.0-1 and 2.0-2,

was the spacecraft pilot for this flight.

The MA-9 space vehicle was launched from the Missile Test Annex at

Cape Canaveral, Florida, at 08:OU a.m. e.s.t. on May 15, 1963 The

flight ended as planned, after completing nearly 22 orbital passes around

the earth, with a successful landing approximately 70 nautical miles

southeast of Midway Island in the Pacific Ocean at 06:2 p.m. e.s.t. on

May l6, 1963. Ground tracks for the 22 orbital passes of the MA-9 space-craft are shown in figure 2.0-3-

The MA-9 mission was a continuation of a pioneering program to acquire

operational experience and information for extended manned orbital space

flight. The objectives of the flight were to evaluate the effects on the

astronaut of approximately 1 day in orbital flight; to verify that man

can function for an extended period in space as a primary operating system

of the spacecraft; to evaluate in a manned 1-day mission the combined

performance of the astronaut and a Mercury spacecraft specifically modi-

fied for the mission; to obtain the astronaut's evaluation of the opera-

tional suitability of the spacecraft and supporting elements for extended

manned orbital flight; and to assess the effectiveness of the Mercury

Worldwide Network and mission support forces during an extended manned

orbital flight. Each of these objectives was satisfactorily fulfilled.

A preliminary analysis of the significant flight data has been made,

and the results are presented in this report. Brief descriptions of the

mission, the spacecraft, and the launch vehicle are followed by the per-

formance analyses and supporting data. All major events of the MA-9

mission, beginning with delivery of the spacecraft to the launch site and

continuing through recovery and postflight examination, are documented.

The graphical information presented herein has been included to

support and clarify the text; however, the reader is referred to Part II,

Data, for a complete presentation, without analysis, of all MA-9 time-

history flight data. Part III, Mission Transcripts, presents essentially

unedited transcripts of the flight communications, the pilot's postflight

self-debriefing, and the formal technical debriefing conducted onboard

the recovery aircraft carrier.

W T O B M T I A I i



Figure 2.0-1.- MA-9 astronaut prior to entering the transferva n i e . a . Q . b ] e launch complex.9Ji«JAep.j ak%.,kQp.3b]ael



Page

Figure 2.0-2.- MA-9 astronaut on the launch-pad gantry prior to fligh



Page 2-4

Latitude, deg

C O M r i D C M T I A f e .





3 - 2

6. A "rate indicate" switch with an automatic and manual

position was added. In the mnaual position, telemetry and

pilot-indicated rates were given continuously; and in the auto-

matic position, the rate indications were cut off from the time

of spacecraft separation plus 5 minutes to 10 minutes before

retrosequence time.

#7- An "out-of-orbit mode" warning light and tone switch

were added. This circuit was available from 5 minutes after

spacecraft separation until the beginning of retrosequence.

3.1.1.2 Reaction control system:

1. The 1-pound and 6-pound thrusters of the reaction

control system (RCS)were replaced with units of an improved

design.

^2. The nitrogen tank in the automatic RCS was pressurizedto 2,800 psi instead of 2,250 psi, as in previous missions.

3. A special corrosion-deterrent paint was applied to

the outside of all hydrogen peroxide (H 0 ) tanks.

U. A dual indicator was added to the instrument panel

to display to the astronaut the regulated nitrogen pressures

in the automatic and manual reaction control system.

5 - The time-delay relay used in the jettison of HO was

changed to extend the jettison time from 60 seconds to

150 seconds.

61. The wall thickness of the expulsion tubes of the

automatic and manual HO tanks was increased from 0.062 inch

to 0.125 inch.

•"-7. A 15-pound capacity HO tank was added in parallel

with the automatic HO tank.

-«8. A manually operated interconnect valve was added to

provide the capability to transfer fuel between the automatic

and manual HO fuel systems.

9. A drain and purge valve was added to the automatic

and manual HO systems.

10. The nitrogen and hydrogen peroxide relief valves were

replaced with units of a more reliable type.



Page 3-3

3.1.2 Life support system. -

3-1.2.1 Environmental control system:

1. The electrical inverters were cooled by using heat

sinks and ducted cabin air circulation, rather than the pre-viously used coolant circuit and heat exchanger.

*2. The CO adsorption capacity was increased by adding

0.8 pound of lithium hydroxide to the CO canister. The amount

of charcoal in the canister was reduced to 0.2 pound.

3. A provision was added for manually sealing the cabin

pressure-relief valve from water leakage at landing.

k. The warning light for indicat ing exce ss water or low

temperature in the suit or cabin heat exchangers was made de-pendent upon the associated dome temperature.

- : ; - 5 - The cabin oxygen-partial-pressure indicator was re-

placed with a dual indicator displaying cabin 0 partial

pressure and suit-circuit CO partial pressure. Also, a warn-

ing light and a tone switch were ad ded to the panel to indicate

excessive CO .

6. A redundant coolant-control valve was added in parallel

with the existing valve for the suit cooling circuit. Also,

these valves and the cabin coolant-control valve were of animproved design.

#7- An oxygen bottle containing h pounds of oxygen was

added in parallel to the primary oxygen bott le.

•;:'8. A 9-pound coolant water tank was added in parallel

with the existing 39-P°und coolant water tank.

9. The suit and cabin freon orifices and check valves

were replaced with units of an improved design.

10. The primary and secondary oxygen high-pressure regu-lators were replaced with modifie d units of an improved design.

*L1. The secondary oxygen-supply system had a warning

light and tone switch for an indicati on to the pilot when

pressure in the oxygen bottle dropped below 6,500 psi.




Page 3 - k

12. The emergency oxygen rate valve was replaced with one

of an improved design.

IJ. The pressurization system for the coolant-water

supply was provided with gas pressure from the suit circuit

only.

Ik. The pressure-sensing circuit that was used to cut

off the ca~bin fan and energize a warning light indicating low

cabin pressure was removed.

-"-15. A timer circuit was installed to actuate the water-

separator sponge for a JO-second period every 10 minutes.

16. A "sponge-squeeze" switch and water-separator travel

indicator lights were added to the main instrument panel to

allow automatic or manual initiation of the water separator

and to provide a visual aid for monitoring the separator piston

position.

17. The absolute-pressure relief valve was removed from

the coolant-water pressurization system.

18. Insulation was added to the CO absorber, and a

deflector was added to the cabin fan outlet to prevent cold

air from the cabin heat exchanger from impinging on the CO

absorber.

19. A screen was added over the relief port of the

negative pressure relief valve to prevent objects from fallinginto the valve opening.

20. The suit inlet for emergency oxygen flow was re-

positioned upstream of the suit-circuit CO partial pressure

sensor. This change permitted purging of the sensor with

100-percent oxygen to verify its operation.

#21. A condensate trap was installed in the suit circuit

to aid in removing water.

22. The suit-circuit ducting, from the water separatorto the junction of the suit-inlet flexible hose, was insulated

to reduce heat loss.

3.1.2.2 Food, water, and waste management:

*1. Condensate and urine transfer systems were installed.



Page 3-5

2. The nylon drinking-tut)e assembly from the 39-pound

coolant water tank was removed.

#3- An expulsion-type drinking-water container was added.

The tank contained a transfer fitting to provide the capabil-

ity of transferring condensate to the tank after the drinkingwater was consumed.

3.1.2.3 Pull-pressure suit:

1. The sealing technique for the faceplate on the helmet

was changed from pneumatic to mechanical. In addition, small

velcro tape patches for attaching the oral temperature pro"be

to the right earcup, radiation film badges attached to the

helmet shell, and an improved helmet tiedown system were in-

stalled.

2. The torso section of the pressure suit was modifiedin the shoulder and wrist areas. In addition, the boots were

made a permanent part of the torso assembly, and they were

provided with improved ventilation.

3 - A poppet-type valve was added to the suit-inlet venti-

lation fitting to prevent water from entering the suit should

the astronaut leave the spacecraft and enter the water after

landing.

h. An additional locking feature was added at the

glove-to-torso connection to prevent accidental disengagement.

•"•5. A urine transfer fitting was added to the suit.

6. The lifevest pack was moved to the front of the

lower left leg since this was a m©re convenient location.

7- Items carried in pockets on the suit included a

handkerchief, pocket folding knife, biomedical injectors,

heavy-duty scissors, and a mechanical pencil.

3.1.2. Personal equipment:

1. The navigation yaw reticle was deleted.

2. An opaque window cover was added to the spacecraft

window, and the red window filter and the map case were removed.



3- A work-table and storage space assembly was added in

front of the main instrument panel pedestal.

U. The flashlight, the food container, the knife mounted

on the spacecraft structure, and the waste container were

removed.

*5- An exercising device of the type used during the

MA-6 flight was added.

6. The rear-view mirror was removed.

3.1.3 Communications system.-

1. The hardline cable which allowed operation of the

voice communications system outside of the spacecraft was

deleted.

2. Nitrogen gas pressure from the manual RCS was used

to deploy the HP recovery antenna.

3- The fingers were removed from the bicone antenna to

improve communications while in the attitude-free drifting-

flight mode.

*k. The back-up UHF voice transmitter-receiver was

removed.

5 - The capability for ground command of the telemetry-

system and radar-beacon operation was incorporated.

6. The HP voice transmitter was disabled from the time

of antenna-fairing separation to the time of HF recovery

antenna deployment to prevent damage to the voice system.

7. A switch was added to allow the pilot to disable the

UHF power amplifier for increased reliability of the UHF trans-

mitter-receiver system.

8. A slow-scan television system was added for real-time

observation of the pilot and spacecraft environment. The TV

transmitter was also available as a backup to the T^VI trans-mitter.

9- The phase-shifter switch was removed, and the phase

shifter was controlled by a relay such that it was on whenever

the C-band beacon was on. Phase-shifter power was changed from

the ASCS bus to the fans bus.



Page 3-7

10. Deployment of the HF orMtal antenna was effected by

only one bellows motor, 10 seconds after spacecraft separation.

11. A control unit was added to allow the pilot to set

a high volume on his receiver with a minimum of side tone

while transmitting.

12. The miniaturized helmet microphone was modified to

include noise rejection characteristics. The helmet was also

modified to include two miniaturized earphones in each ear-

piece.

J.lA Mechanical and -pyrotechnic systems. -

*1. The periscope was removed, and a spring-loaded device

was used to close umbilical door.

2. The SOFAR bomb on the main parachute riser was setto detonate at a depth of 3,000 feet.

3. The main and reserve parachute deployment bags were

modified for increased reliability.

U. The landing-bag release system was modified to im-

prove its reliability.

5. A redesigned survival kit pan was installed.

6. One of the two squibs in the retropackage explosive

jettison bolt was disarmed because it contained a ground loopcircuit.

7 - The explosive-actuated hatch actuator cap was vented.

3.1.5 Electrical and sequential systems.-

1. A redundant 3-volt power supply was added for instru-

mentation reference. It could be actuated by the astronaut for

an instrumentation reference of flight-critical items in case

of failure of the primary 3-volt power supply.

2. The eight day correlation clock was removed.

3- The satellite clock was powered by the 2 v d-c main

bus rather than the isolated bus.

^. An "off" position was added to the instrument-panel

warning-light circuit to turn off all lights except those for

the satellite clock for improved pilot dark adaptation.

iV711r UL/IMl 1lALi



p a g e 3 - 8 . OOHriDCNTIAL

5. The main inverters were replaced with those of an im-

proved design having superior thermal characteristics and a

greater efficiency.

#6. In order to provide increased power for the longer

MA-9 mission, five 3 000 watt-hour and one 1,500 watt-hourbatteries were flown in spacecraft 20 instead of the three

3,000 watt-hour and the three 1,500 watt-hour batteries flown

in spacecraft l6 (NLA-8).

7- The fire retro, green telelite was made dependent on

the ignition of all three retrorockets rather than being

dependent on the ignition of the third retrorocket, as was the

case for spacecraft 16.

8. A switch was added to provide the astronaut with the

capability of turning off the flashing recovery light to con-

serve power during daylight hours.

9. An auxiliary portable light was installed for the

astronaut's use during flight.

10. A switch was added to allow turning the prelanding

buses back on once the landing relay timed out to permit post-

landing blood-pressure and EGG recordings.

11. A tower-jettison arm relay was added in the tower-

jettison circuit to prevent inadvertent firing of the tower-

jettison rocket.

12. A green telelite was provided to give the astronaut

an indication of umbilical door closure, and a red telelite

was provided to indicate when the door was open. The door posi-

tion was also monitored by telemetry.

13. A tower-separation abort-interlock relay was added on

the 2h- v d-c isolated squib bus. If an abort signal were to be

received by the spacecraft and power from the main squib bus

were lost, initiation of retropackage jettison would occur when

spacecraft separation was sensed. Tower-jettison rocket igni-

tion would occur when tower-clamp-ring separation was sensed.

1 . Both the main and isolated bus circuits for the retro-

rocket ignition squibs were controlled by the retrofire arm

circuit. This change was made to improve the switching arrange-

ment for the astronaut. Only the main squib-bus circuits were

so controlled on spacecraft 16.



Page 3 - 9 * " -;

15. The automatic retrofire arm feature was made functional

for all retrosequence modes except manual retrofire, with or

without attitude permission. As a result,, whenever the pilot

elected to ignite the retrorockets using the retrofire switch,

the retrorocket arm switch had to "be in manual position.

16. The Mayday circuit was powered from the main bus; it

was powered from the main squib bus on spacecraft 16.

IT- A rescue aids switch was rewired to provide the

astronaut with the capability of manually extending the HF

recovery antenna after landing.

18. Three diodes were put in series with the flashing

light to prevent damage to the light, because the light was

powered by the 6-volt isolated bus, rather than its own

self-contained battery.

19. A resistor was added to the command input circuit of

the programer to reduce the programer's susceptibility to

transient voltage spikes.

20. A standby inverter automatic tone generator was added

to indicate automatic switching of the inverter to either a-c

bus.

21. The emergency reserve parachute deployment and the

emergency landing-bag-deployment circuits were powered through

switch fuses.

22. The satellite clock, pilot, and ground command retro-

sequence signals were powered through a common fuse switch.

23. An emergency spacecraft separation bolt relay was

added in the spacecraft-separation pull-ring circuitry to allow

the pilot to fire the escape rocket with isolated bus power

only.

2 -. Two spacecraft-separation-sensor relays were added to

the isolated squib bus. One relay improved the reliability of

the maximum-altitude sensor and the other was used in the

tower-separation abort-interlock relay circuitry.

25- Automatic 21,000-foot drogue-parachute deployment was

made more reliable by paralleling the main and isolated squib

arming circuits.



Page 3-10

3.1.6 Instrumentation system.-

1. A switch was added to provide the astronaut with the

capability to remove electrical power from the R- and

Z-calibration relays. This switch would be used to stop

calibrations should the programer fail 1;operform this function

as planned.

2. The HF telemetry system was deleted from the spacecraft.

3. The low-level commutator and temperature-survey pick-

ups were removed from the spacecraft.

^ 4 - . Several timing functions were added to the programer.

In spacecraft 1.6, its only function was programing the water-

squeezer operation.

5. The frequency of the voltage-controlled oscillator

for the automatic-solenoid malfunction detector was changed to

3-9 kc to provide better tape-recording reproduction of this

function.

6. The "A" package of the instrumentation system was

modified to protect the d-c amplifiers from over voltage and

to provide for a better reading of the 115 v a-c fans bus.

7- The oxygen-quantity indicator was expressed in per-

cent. Maximum indication was 250 percent on primary and

125 percent on secondary.

8. The R- and Z-calibration signals were no longer

initiated only on ground command. They were initiated when

the telemetry transmitter was energized through the command

link and when the tape recorder was programed.

9. The tape-recorder operation was no longer completely

continuous. It was modified to run at a speed of rr inch

per second, giving a longer recording capability, and its

operation was either off, programed, or continuous, as selected

by the pilot.

10. A three-position switch with off, continuous, and

ground-command positions was added for astronaut control of

telemetry transmitter operation.

11. The astronaut-observer camera was replaced with a

self-contained, hand-held moving-picture camera. This camera

was a l6-mm type that could be mounted on the instrument panel

for observation of the pilot or on a special bracket to photo-

graph through the window.



12. The retrorocket temperature sensor was moved from

the right to the bottom retrorocket.

13. The low-fuel-pressure -warning switch was changed

from a 1,580-psi switch to a 2,200-psi switch, and the auto-

matic ECS high-pressure transducer range was 0 to 3,500 psi.

1 . The cabin and heat exchanger dome temperature light

and tone alarm replaced the excess cabin and suit excess

water light and tone alarm.

15. Suit outlet temperature was sensed by the body tem-

perature sensor when the sensor was not measuring body

temperature. Temperature range of the sensor was 75° F to

108° F.

16. Temperature sensors were installed in all three fuel

tanks, and the temperatures were displayed on an instrument-panel indicator.

17. Standby inverter and cabin heat-exchanger outlet

temperatures were telemetered to the ground.

18. The blood-pressure-measuring system controller was

changed to provide easier gain adjustment.

3.1.7 Heat-protection system.-

1. Six bolts were added to the ablation heat shield to

retain the shingle portion of the shield in case of delamina-tion at landing.

2. In addition to the previously flown rectangular paint

patch, two new types of paint were added for evaluation.

3.1.8 Experiments.-

1. A tethered balloon similar to that flown during the

MA-7 mission was packaged in the antenna canister.

2. A self-contained flashing beacon was installed on the

retropackage.

3- Two geiger counters, a dosimeter, four film badges,

and a Schaeffer radiation package were installed on or in the

spacecraft to determine radiation exposure during the flight.



Page 3-12

U. A special 35-™m camera was added for use in photo-graphing the zodiacal light and the airglow layer.

5. A special 70-rnm Hasselblad camera was added for usein taking general color photographs, infrared weather photo-

graphs, and horizon definition photographs.



Page 3-13

TABLE 3.1-1.- SUMMARY OF SPACECRAFT SYSTEM MODIFICATIONS

System

Spacecraft control system:

Automatic stabilization and control system

Reaction control system

Life support system:

Environmental control system

Food, water, and waste management systems

Full-pressure suit

Personal equipment

Communications systems

Mechanical and pyrotechnic systems

Electrical and sequential systems

Instrumentation system:

Telemetry and sensor systems

Instrument panels and consoles

Heat protection system

Scientific experiments

Total

Number of changes

(a)

10

19

31776

15

8

35

16

5^

2

5

215

The total is not an accurate indication of the total number of

spacecraft modifications, since the nature of certain system interfaces

requires that some changes be repeated for more than one system. In

addition, because of this repetition in some areas and the fact that

some modifications which are counted only once can reasonably apply

to another system, the numbers in the column should be regarded as

approximate and only a gross indication of the degree of modification

for any given system.



Page 3 -

TABLE 3.1-2.- WEIGHT AND BALANCE DATA FOR SPACECRAFT 20

Parameter

Weight, Ib . . . .

Center -of -gravitystation along:

X-axis, in. . .

Y-axis, in. . .

Z-axis, in. . .

Moments of inertiaaround :

Roll axis, I ,

2Z

slug-ft ...

Pitch axis , I ,* V

2slug-ft . . .

Yaw axis , I ,

slug-ft2. . .

Lift-off

\, 330. 82

-0.13

-0.28

167.81

365.8

7,900.6

7,90 .0

Orbital

phase

3,033-35

-0.21

-0.1U

120.82

298.9

653.0

656.8

Reentry

phase

2,681. U5

-0.20

-0.15

12 .68

280.9

563.0

571-3

At mainparachute

deployment

2,563.89

-0.20

-0.15

122 . 22

27 .8

^

505.5

Postlanding

2 , -00. 3

-0.49

-0.06

119.63

269.5

1+38.7

W8.1

11/H7



3-2 Launch-Vehicle Description

Page 3-15

The MA-9 launch vehicle, the Atlas 130D, vas an Atlas series D

missile modified for the mission as on previous Mercury-Atlas flights.

A general launch- vehicle description may "befound in the NASA Project

Mercury Working Paper No. 223A, "Manned One-Day Mission Mercury Spacecraft

Specification Document. "

The MA-9 launch vehicle was very similar to the one used for the

MA-8 mission, and only necessary changes were made. The following is a

summary of the detailed configuration changes from the MA-8 launch

vehicle, the Atlas 11JD.

(1) A plastic liner was incorporated in the wear-ring area of the

turbopump to guard against a possible failure resulting from excessive

rubbing.

(2) A temperature sensor was added to the head- suppress ion valve,

which is located on the sustainer engine housing, to determine the temper-

ature of the head suppression valve during flight and required the

installation of three wires to the harness, a resistor, and two wires to

the telemetry package.

(3) The clips, which are used to attach the shroud to the forging

that holds the yaw activator, were modified so that the clips would not

ride the radius of the vernier engine gimbal shaft.

The mount for the secondary range-safety command battery was

redesigned to reduce weight and provide greater ease in manufacturing.

(5) A redundant circuit, including instrumentation, was provided

in the engine relay box to improve the reliability of the sustainer-

ignition-stage control-valve circuit.

(6) The power pickoff point for the telemetry and instrumentation

system was changed from the power plug to the changeover switch. This

change provided the telemetry system with a 115-volt (a-c) -00-cycle

instrumentation point which would not interfere with the guidance system.

(7) The lox overfill probe was relocated and redesignated "Sequence

II Level Probe." This modification provided for a repeatable method of

determining the proper level at ignition start by maintaining the lox

level at this probe.

(8) A printed circuit board in the programer canister was redesigned

to remove the possibility of a locating pin's shorting a transistor on the

circuit board.



P a g e 3 - 16 C O N F I D E N T I A L

(9) The event times for the flight programer were changed to be

compatible vith the staging discretes of the latest launch trajectory.

(a) BECO - 132.9 sec instead of 129-9 sec

(b) Sustainer pitch program duration - 13-5 sec instead

of 16. U sec

(c) Initiation of guidance after BECO - 22.5 sec instead

of 2 .0 sec

(d) SECO - 303.6 sec instead of 305.1 sec

(10) To reduce the possibility of an undesirable lift-off clockwise

roll transient, the booster-engine yaw actuators were offset (see

fig. 3-2-1) as follows to produce a counter-clockwise roll moment:

Booster-engine no. 1 - yaw actuator lengthened by 0.0 ± 0.02inch

Booster-engine no. 2 - yaw actuator shortened by 0.0 ± 0.02

inch

The offsets were checked by the usual level method and by using a new

alinement jig supplied by the engine contractor.

(11) The temperature sensor in the sustainer-engine lubricant tank

was relocated to the aft 20 percent of the tank to provide a temperature

study of the lubricant as it is consumed.

(12) A redundant path to ground was provided for the shielding in

the autopilot harness.

(13) The boat cover in the sustainer engine area was restrained by

a spring which had a tensile strength approximately twice that of the

spring used in the MA-8 launch vehicle. This change was made to provide

better thermal protection of wiring harnesses.

A preflight purge of the boattail area with 100-percent gaseous.

nitrogen was incorporated to reduce the possibility of fire.

(15) The propellent utilization (PU) manometer was calibrated for theAtlas-D tank, rather than for the Atlas-C configuration.

(16) A microswitch which indicates full lox-valve travel was rewired

to permit inclusion in the ignition-stage sequence circuit to reduce the

possibility of a lox-pump failure.

(17) The wiring technique for the autopilot in the flight control

section was modified to improve its overall reliability. These units were

replaced at the factory prior to delivery of the launch vehicle to the launch

site.



SOMflDCNTIALPage

Figure 3.0-1.- MA-9- MA-9 space vehicle prior to launch.



C O N F I D E N T I A LPage 3-

Figure3- 2 K,JVW -.W y off configuration.



3 M E I P E N T I A L Page 3

Figure 3.1-1.- MA-9 spacecraft and adapter prior to lift-off,



C O N F I D E N T I A LPage3-20

Previous position

Offset position



Page k - 1

k.O TRAJECTORY AND MISSION EVENTS

Sequence of Flight Events

The times at which the major events of the MA-9 mission occurred

are given in table U.l-1.



Page h - 2

TABLE IK 1-1.- SEQUENCE OF EVENTS

Event

Q

ELanned time ,hr :min : sec

Actual time,hr:min:sec

Difference ,

sec

Launch phase

Booster engine cut-off (BECO)Tower releaseEscape rocket ignitionSustainer-engine cut-off (SECO)discrete signal

Tail -off complete

00:02:12.800:02:35.600:02:35.6

(b)00:05:0 .5

00:02:12. !< •

00:02:36.300:02:36.3

00:05:03.000:05:03.2

-OA

0.70 . 7 .

-1.3

Orbital phase

S/pacecraft separationRetrofire sequence initiationRetrorocket no. 1 (left)Retrorocket no. 2 (bottom)Retrorocket no. 3 (right)Retrorocket assembly jettison

00:05:05.5

33:58:5 .133:59:25.133:59:30.1

33:59:35.13*1:00:25.1

00:05:05.3

33:58:5933:59:30

33:59:3533:59:J>k:OQ:hk

-0.2

.9

.9.9

c^-9C

l8.9

Reentry phase

Communications blackout0.05g relay actuation

Drogue parachute deploymentMain parachute deploymentSpacecraft landing

3 :07:56.13 :08:36.1

3 :13:53.13 :15:21.13 :19:56.1

3lv:08:17(e)

3 :1 :033 :15:333li:19:lj.9

20.9d(l)

n.99d(-9)

(g )

Preflight calculated, based on nominal Atlas performance.

Planned trajectory times are based on tail-off-complete conditions, ratherthan SECO conditions.

Ketrorocket assembly jettisoned manually.

Difference between the actual and the postflight-calculated reentry eventtimes, shown in parentheses, is based on actual insertion parameters.

eQ.05g sequence disabled prior to retrofire.

fDrogue parachute deployment initiated manually.

^Landing time could not be established accurately.



Page h - 3

k-,2 Flight Trajectory

The trajectory for the MA-9 flight is discussed in three conven-

ient phases: launch, orbital, and reentry. In all trajectory figures,

the trajectories marked "planned" are preflight-calculated nominal tra-

jectories and the trajectories marked "actual" are based on Mercury net-

work tracking data. The altitude-longitude profile for the entire flight

is presented in figure U.2-1.

A comparison of the planned and actual trajectory parameters is

given in table .2-1. The differences between the planned and actual

flight trajectory parameters are a result of the actual cut-off conditions

being slightly different than the planned conditions and the atmospheric

density profile on the day of the actual flight being different from that

assumed for the preflight-calculated trajectories.

The launch trajectory data shown in figure ^-.2-2 are based on the

real-time output of the Range Safety Impact Predictor Computer (lP-709 -),

which used Azusa MK II and Cape Canaveral FPS-16 radars, and the General

Electric-Burroughs (launch-vehicle guidance) computer. The data from

these tracking facilities were used during the time periods listed in

the following table:

Facility

Cape Canaveral FPS-16

Azusa MK II

General Electric -Burroughs

Elapsed time, minrsec

.0 to 00: 36

00:36 to 01:0

01:0 to 05:55

The orbital portion of the trajectory, shown in figure k.2-3, was

derived by starting with the spacecraft position and velocity vector

obtained at the beginning of the second pass over Bermuda, as deter-

mined by the Goddard computer using Mercury network tracking data.

The Bermuda vector was integrated backward along the flight trajectory

to orbital insertion and forward to the time of the Cape Canaveral vector

at the end of the l8th pass. The Cape Canaveral vector was then integrated

forward to the start of retrorocket ignition in the 22nd pass. These inte-

grated values were in good agreement with the values measured by the launch-

vehicle guidance system at orbital insertion. They were also in good agree-ment with the position and velocity vectors determined by the Goddard com-

puter for passes near Eglin Air Force Base, Florida (end of 3rd pass),

Eglin Air Force Base (updated at end of 13th pass), and Cape Canaveral

(end of l8th pass); thus the validity of the integrated orbital portion

of the flight trajectory was established.

C O N F I D E N T I A L "



Page

The orbital lifetime of the spacecraft, bas.ed on the 1959 ARDCatmosphere, was calculated to be 92 passes. After nearly 22 passes,the decay in apogee was 7-1 nautical miles and the decay in perigeewas 1.6 nautical miles.

The reentry portion of the trajectory, shown in figure 4-.2-4, wasderived by starting with the spacecraft position and velocity vector,as determined by the Goddard computer, obtained at the end of thel8th orbital pass near Cape Canaveral, Florida. Integrating forwardalong the flight path to retrorocket ignition and,after introducingnominal retrofire conditions, continuing the integration through space-•craft landing yielded the reentry trajectory. Nominal retrofire condi-tions include a retrorocket total impulse of 38,975 l^-sec at spacecraftattitudes of -34° in pitch and 0° in roll and yaw. The spacecraft weightat retrofire was estimated to be 2,979 pounds by using data obtained fromthe. Mercury network stations. The times of communications blackout andmain-parachute deployment from the integrated reentry trajectory were ingood agreement with data from the Mercury network stations and the space-craft onboard measurements. In addition, the landing point from theintegrated trajectory was in good agreement with the retrieval pointreported by the recovery ship. The agreement in these events serve toconfirm the validity of the integrated reentry portion of the flighttrajectory.



Page k - 5

TABLE IK 2-1.- COMPARISON OF PLANNED AND ACTUAL TRAJECTORY PARAMETERS

Condition Planned Actual Difference

Cut-off conditions (including tail-off)

Range time min°secGeodetic latitude , deg North . . .

Longitude deg WestAltitude feet

Altitude nautical miles

Space -fixed velocity, ft/sec . . .

Space -fixed flight -path

Space -fixed heading angle,deg East of North

3014,505:0*4-.5

30.432372.5023

528,l4-02

87.0

437.7

25,715.3

0.0016

77.4909

303.2

05:03.2

30.1t857

72.5178

529,73587.2

437.725,714.0

O . O Q l i - 7

77-5510

-1-3-00:01.3

0.053U

0.0155

1,3330.2

0

-1.3

0.0031

0.0601

Postposigrade firing conditions

Range time sec

Range time mint sec ........Geodetic latitude, deg North . . .Longitude deg West

Altitude feetAltitude nautical miles

Range, nautical milesSpace-fixed velocity, ft/sec . . .

Space-fixed flight-path angle,des ..

Space-fixed heading angle, degEast of North

306.505:06.530.11-621

72.3552528,434

87.0

445.525,736.3

-0.0014

77.5695

306.3

05:06.3

30.531572.2897

529,79387.2

1+49.825,734.9

0.0023

77.6731

-0.2

-00:00.2

0.0694

-0.0655

1,3590.2

4.3.-1.14-

0.0037

0.1036

Orbital parameters

Perigee altitude, statute miles . .

Perigee altitude, nautical miles

Apogee altitude , statute miles . .

Apogee altitude, nautical miles . .Period, min:sec

Inclination angle deg

100.1

87.0165.7

i44.o88:1(432.52

100, 3

87.2165.9144.288:45

32.55

0.2

0.2

0.2

0.200:01

0.03





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5 -

5. 0 SPACECRAFT P E R F O R M A N C E

The spacecraft as an entity performed adequately. Some system

anomalies were experienced, and analyses of these are discussed in the

following paragraphs. Also discussed, from an overall mission viewpoint,are the spacecraft systems' general performance. In addition, a descrip-

tion of each spacecraft system is presented. This description is pre-

sented in terms of the major changes made since the MA-8 mission, and

reference should be made to section 3-1, Spacecraft Description, for a

listing of all significant spacecraft changes.

5-1 Spacecraft Control System

All spacecraft control system components functioned normally until

approximately 28:3 :3 .> at which time the 0.05g relay circuit was actuated.

The astronaut did not report this event on the onboard tape until approxi-mately 28:59-'00. The .reason for this delay in reporting was twofold:

1. The control mode at the time of 0. 05g relay activation was manual

proportional with gyros caged, and

2. The spacecraft warning lights switch was in the off position

because the astronaut was engaged in taking photographs.

When the warning lights switch was placed in the dim position, the

0.05g green light was noted. From 28:3 :3 until the end of the mission,

the amplifier-calibrator (amp-cal) was locked in the 0.05g configuration.

Operation of the automatic stabilization and control system in this con-figuration resulted in damping about the pitch and yaw axes and in a

-12°/sec rate command in the roll axis unless one of the manual control

modes was selected.

5-1.1 System description.- The spacecraft control system is designed

to provide stabilization and orientation of the spacecraft from

immediately following spacecraftlaunch-vehicle separation until

deployment of the main parachute. The system is capable of

operation in the following modes:

1. Automatic stabilization and control system (ASCS) with

alternatives of orientation, orbit, and auxiliary damping modes.

2. Fly-by-wire (FB¥), which is an electrical "on-off"

command of the automatic reaction control system (RCS) thrusters

initiated by the astronaut's control stick. Astronaut's choice

of high and low thrusters or low thrusters only (FEW low) is

available.

3. Manual proportional (MP), which is a mechanical command

of the manual RCS thrusters initiated by the astronaut's control

stick.

V^Vr.1T.TJI-LrEjl M.'M.£ L ^ ^



p a g e 5 - 2

Modes 1 and 2 utilize the automatic ECS fuel supply, while mode3 utilizes the manual RCS fuel supply. Combinations of modes1 and 3 or 2 and 3-are available to provide "double authority"at the astronaut's discretion.

Major changes made to the control system since the previousmission are as follows: the vertical attitude gyro was modifiedto provide the capability to cage and uncage at -3 °, whichsimplified the astronaut's task in alining the gyros to retro-attitude (orbit attitude); the rate stabilization and controlsystem (RSCS) was removed to reduce weight; the ASCS rate gyroswere rewired to run continuously when the ASCS was powered upuntil antenna fairing separation; the ASCS mode switch was changedto provide a means of deenergizing the automatic RCS solenoidcircuitry; a horizon-scanner power circuit was incorporated intothe attitude gyro switch, which powered up the scanners only whenin the "slave" position; and a JO-second time delay was added to

the horizon-scanner slaving circuitry to allow time for scannerwarm-up.

5.1.2 Performance analysis.- The performance of the spacecraft controlsystem was completely satisfactory during the first l8 orbitalpasses. Outputs of the gyros and the horizon scanners agreed towithin 2° during the periods of scanner slaving.

The ASCS orbit mode limit cycles in the pitch, roll, and yaw axeswere relatively balanced; that is, there were an equal number ofpulses on both sides of attitude gyro null. For the greaterpor-tion of operation in the ASCS mode, the yaw-axis limit cycle

reached +10° in periods of less than 3 minutes. However, moreactivity in this axis was expected because of the cross-couplingeffect between the roll and yaw slaving circuitry. The averagelimits of the roll axis cycle were ±8.5% and the limits of thepitch axis cycle were +7°. As was noted during the MA-8mission,pulse durations were not sufficient to limit the orbit cycle towithin the more desirable ±5.5°- However, this condition is notconsidered to be detrimental and does not appreciably increasefuel consumption.

At 28:3 :3 ; the 0.05g relay circuit actuated and locked in. Atapproximately 29:^9tOO? "the astronaut powered up the ASCS bus and

verified that the amplifier-calibrator was in the 0.05g configu-ration by noting that the attitude indicators would not respondto spacecraft attitude changes. He then made a reentry roll ratecheck at about 31:17'-00 and verified that the ASCS would functionnormally in the 0.05g logic circuit during reentry. However,planned use of the ASCS for reentry was abandoned at about 33:07:00

when it became evident that the ASCS bus was not receiving powerfrom either the ASCS a-c main inverter or the standby a-c inverter.

CONFIDi





p a g e 5 - * O O M T D D P i T I A L tner*

The shorting and charring of the power plug and the 0.05g relay •circuit actuation occurred independently; however, both mal-functions apparently resulted from the presence of moisture.

5-1.^ Control system -utilization.- Immediately after spacecraft

separation at 00:05:0^, the astronaut switched to the auxiliarydamping mode for k seconds. At 00:05:1? he began to executethe yaw turnaround maneuver by using the fly-by-wire lowthrusters. By 00:08:33> the turnaround maneuver was completedand the control system was placed in the ASCS orbit mode. Thismode was employed for approximately 15 percent of the time thatspacecraft power was utilized. Fly-by-wire low was used forextended maneuvers and experiments almost exclusively through-out the mission; and, therefore, a minimum usage of the auto-matic RCS fuel supply resulted. Only one automatic high thrusterwas actuated during the orbital phase of the mission, except fort h e momentary high thruster action in the auxiliary damping mode

after spacecraft separation. The negative roll high thrust unitwas utilized prior to retrofire during the epecial 0.05g testat about 31:17:00.

Attitudes for retrofire were maintained by using the manualproportional mode, with fly-by-wire high and low ready as abackup. The astronaut maintained the spacecraft attitudesextremely well during the retrofire period, as was evidencedby the spacecraft's proximity to the planned landing position.Control during reentry was maintained by using the manualproportional and fly-by-wire modes simultaneously. The maximumthrust of ^9 pounds about the pitch and yaw axes was used by

the astronaut in maintaining control during reentry.

5.1.5 Reaction control system.- The major changes made to thereaction control system (RCS) since the previous missioninclude the addition of a 15-pound capacity hydrogen-peroxidetank (see fig. 5-1-5-1) in parallel to the automatic system,the removal of poppets and springs from the check valve atthe outlet of each fuel tank, the incorporation of an RCS inter-connect valve between the automatic and manual fuel system, theinstallation of a main instrument panel indicator to enable theastronaut to monitor regulated nitrogen pressure for both thefuel systems, and an increase in the wall thickness of the

expulsion tubes for the fuel tanks from 0.062 inch to 0.125 inch.

An analysis of the onboard data confirmed the astronaut's reportof satisfactory performance of the RCS throughout the flight.

One instance of thruster "tail-off" in the yaw-right manualthruster was noted by the astronaut. The incident occurredapproximately 13 hours after lift-off during a flight period



-fiDMflDENTIAL -

for which no onboard data are available. A postlaunch

evaluation of this minor anomaly indicates no reason for

this action.

The astronaut reported that he had intended to use FEW for

reentry but that the FEW pitch-up high thruster was slowto light off, and therefore, he selected manual proportional.

An analysis of the onboard data indicates that the FEW pitch-

up high-thruster command was maintained only 0.2 second and

that thruster operation for this brief period was normal. This

short pulse duration, in conjunction with the fact that there

had been no previous use of the high thruster to achieve thrust

chamber warm-up, probably reduced the pilot's confidence in

the FEW mode and led him to elect "double authority" control

during reentry.

The amount of fuel used during the mission is shown in

table 5-1-5-1- Approximately 7 pounds of manual fuel and21 pounds of automatic fuel were jettisoned after reentry.

Fuel supply pressure readings varied during the mission because

of cabin temperature changes, with the automatic system trans-

ducer being affected the most.

Fuel jettison was initiated after reentry as planned at approxi-

mately 3 :15:55- The total time required to jettison the re-

maining fuel from the automatic and manual systems was approxi-

mately 1 minute and 33 seconds. The RCS interconnect valve

was utilized to permit the manual system fuel to jettison

through the automatic system pitch and yaw high thrust chambers.

Solenoid-valve inlet temperatures for the 1-pound yaw-left,

the 1-pound pitch-down, and the 1-pound roll-clockwise thrust

chamber assemblies in the automatic RCS were measured during

the flight and are shown in figure 5-1-5-2. The maximum

temperature at the solenoid inlet recorded during the orbital

phase was approximately 103° F. This temperature was measured

for the 1-pound roll-clockwise thruster at 11:20:00. The

minimum temperature recorded was approximately 53° F and was

measured for the 1-pound pitch-down thruster at approximately

22:30:00.







5 - 8

In addition to system changes to accommodate the extended flight,

other modifications from the spacecraft 16 (MA-8) configuration

were required to improve system performance. The major changes

in this category are discussed in the following paragraphs.

Because of the partial blockage problem experience in MA-8, asuit bypass CCV was installed in parallel with the existing CCV

for redundancy.

A condensate trap, shown in figure 5-2.1.1-1, was installed at

the suit-inlet port to extract free condensate water which would

adhere to the inner wall of the water separator and thus escape

extraction. The inner wall of this trap was a wicking material,

which has the property of passing water, but not gas, when the

material is wet.

The cabin-pressure relief valve was equipped with a water sealing

device to enable the astronaut to lock this valve and prevent

sea water from entering the spacecraft after landing.

A sensor for measuring carbon-dioxide partial pressure (PCO ) was

installed in the suit circuit to indicate CO concentration and

actuate a warning tone and light at a PCO value of 8 mm Hg.

The suit inlet for emergency oxygen flow was repositioned upstream

of the PCO sensor, and thereby permitted a purging of the PCO

sensor with 100-percent oxygen to verify its operation.

The suit-circuit, from the water-separator to the junction with

the suit-inlet flexible hose, was insulated to reduce, heat loss;

and the LiOH canister was insulated to minimize condensation.

The dome temperature, which was monitored by the astronaut during

the MA-8 mission, was also monitored on telemetry and recorded

for MA-9.

5.2.1.2 System performance: The analysis of the ECS performance during

the MA-9 mission was dependent upon the results of preflight tests,

real-time telemetry data, the pilot's inflight voice reports, on-

board recorded data, and postflight inspection and test results.

Launch phase: The suit-inlet temperature was 55° F at astro-

naut insertion into the spacecraft and gradually increased to

6l° F during the freon cooling period prior to lift-off. The

oxygen partial pressure (PO ) readout for the cabin was 0.6 psi

below total cabin pressure after the cabin was purged with

oxygen at the launch site. A gas analysis at this time indicated

98-percent 0 at a cabin pressure of 1 .9 psia. At lift-off,





Page 5-10

it is concluded that the C0p rate of 26l cc/min is believable

for the orbital period. Since the faceplate was estimated bythe astronaut to have been open approximately one-third ofthis time, an undetermined amount of CO escaped into the

cabin. This CO would raise the cabin PCO_ reading and leakoverboard. Since there was no measurement of PCO? in the

cabin, it is impossible to estimate from a systems standpointthe quantity of COp that was lost when the faceplate was open.

An estimate of the maximum PCO^ in the cabin based on physio-

logical considerations is contained later in this section.Based on this estimate, a negligible amount of C0p was lost

through the open faceplate. In any case, the figure of 26l cc/minfor the average CO^ production rate is a minimum value.

The cabin-leakage rate determined several days prior to thefirst launch attempt was ^85 cc/min at 19.7 psia and 70° F.This leakage rate was determined by a stabilized flowmetermeasurement. Subsequent to this check the hatch was removedand replaced several times. A brief leakage check on the dayof the launch showed no detectable leakage. This check, however,was quite gross and has no particular significance. A grossleakage-rate determination after launch was obtained from thereading of cabin pressure decay from relief-valve seal-offpressure to the point when cabin-pressure regulation began.This determination showed a leakage of 510 cc/min corrected

to 19.7 psia and 70° F.

Extrapolating the prelaunch leakage determination from the sea-

level condition to the orbital condition shows that the equiv-alent cabin leakage rate during the orbital period would be

0.528 X 10 Ib/min. This leakage rate is determined bycomputing the equivalent orifice area required to leak the^-85 cc/min at sea level and then using this area to computethe choked orifice flow during the orbital period.

Based on the average C0? production rate and a respiratory

quotient (RQ)of 0.83., as noted in a later paragraph, the astro-naut oxygen consumption from 03:05:00 to 3 :05:00 would be1.84 pounds. The leakage for the same period based on extrapo-lated precount measurements amounts to 0.98 pound. The sum ofcomputed astronaut consumption and computed leakage rate istherefore 2.82 pounds as compared with the calculated bottledepletion of 2.18 pounds.





Page 5-12

average cabin temperature during the evaluation was 90° F.At 32:05:00 the astronaut terminated the evaluation and

turned on the cabin cooling system, as planned, in prepara-tion for reentry.. The cooling system responded rapidly, asevidenced by a drop in the cabin-heat-exchanger outlet

temperature.

At 52:15:00, a marked increase occurred in the PCOp level,

which reached 2..8mm Hg by 32:45:00. Prior to 32:15:00, thePCOp was less than 0.1 mm Hg. At 32:44:20, the astronaut

selected the emergency rate mode of 0? flow to purge the PCO

transducer and test its readout validity. This mode was inoperation for 25.5 seconds. The PCO decreased to approxi-

mately 1.9 mm Hg at 32:44:56 thereby verifying the validity ofthe readout. The PCO? again continued to rise gradually until

33:59:30, which is the exact time of ignition for the firstretrorocket. At this time, the PCO? indication decreased

sharply to a negative voltage output. Test experience hasshown that the sensor indication will not normally change asrapidly as it did during this brief period. Postflight cali-bration of the PCOp sensor did not indicate a significant

shift in its calibration.

The LiOH canister has been tested extensively under normalgravity conditions,- and the operational life of the 5-4-pound

charge of LiOH was well established. Prior to astronaut in-sertion on the day of the launch, the effective life of theLiOH canister had been reduced by approximately 8 hours becauseof usage during systems tests conducted for the unsuccessfuland successful launch attempts of May l4 and 15, respectively.Calculations made after the launch attempt indicated that thecanister capability was sufficient to accommodate the mission

with at least a 3-hour margin. Consequently, it was knownthat PCO might build up during the latter part of the mission,

since the useful life of the canister is defined as time re-

quired to reach a PC02 level of 8 mm Hg, rather than the time

to the first indication of C02 buildup. Postflight chemical

analysis of the canisters (table 5.2.11.2-1) indicates that theastronaut's carbon dioxide production was below the design levelof 400 cc/min and that there was 2.18 pounds of LiOH remainingin the canister. However, postflight analysis showed that somechanneling of flow occurred in the flight canister. Thischanneling of flow could explain the indicated PCO buildup.

"CONFIDENTIAL











1ENTIAL Page 5 - 1?

assembly which could be inserted into the desired urine transfer

fitting located on the right-hand console. Pumping the syringe

causes urine to flow from the suit bag to the desired storage

bag beneath the couch. At the completion of urine transfer,

the protected needle is withdrawn from the urine transfer fitting,

and the rubber diaphragm within the fitting automatically seals.

The condensate collection system was capable of storing up to

12.36 pounds of condensate in permanently installed containers

and 6 pounds of condensate in small 1-pound plastic bag con-

tainers, giving a total storage capacity of over 18.36 pounds.

The condensate collection components are the squeezer-type water

separator, a condensate trap, the ^(--pound-capacity condensate

tank, a 3-86-pound-capacity bag installed under the head rest,

and the 4.5-pound potable water tank, which could be utilized

after its contents were depleted. A new vent was extended into

the condensate tank interior as a standpipe. The interior, or

tank side, of the standpipe was coated with a non-wetting

silicone material to prevent water from creeping up the stand-

pipe and out the vent under zero-g conditions. A separate

transfer syringe, similar to that in the urine transfer system,

was mounted on the left side above the astronaut's head. The

inlet of the syringe was connected to the 4-pound-capacity con-

densate tank. This syringe could be used to transfer water

from the 4-pound capacity condensate tank to either the 3-86-pound-

capacity bag or the 4.5-pound~capacity drinking-water tank. The

condensate water was to collect initially in the 4-pound- capacity

condensate tank from both the water separator and the condensate

trap and then be transferred to the 3-86-pound-capacity bag under

the couch headrest. After the 3-86-pound-capacity bag under the

couch was filled, the condensate was then to be transferred to

the 4.5-pound-capacity drinking-water tank, if the drinking

water had been consumed.

After these three containers were filled, the procedure would

then be to transfer additional condensate to 1-pound plastic

bags included in the special equipment storage kit. These

plastic bags could be filled in the same manner as that used

when adding water to the food packets.

5.2.3.2 Systems performance: The 4.5-pound-capacity water tank operated

satisfactorily, except for the drinking-water mouthpiece. The

astronaut reported that the mouthpiece leaked about the valve

body when he opened it to add water to the dehydrated-food bags.

A postflight failure analysis of the valve revealed that if any

back pressure, such as that created when filling the food bags,

is imposed at the valve outlet, it would leak between the valve

locknut and the mouthpiece (fig. 5-2.3-1). This leakage could



Page 5 - 18

Toe minimized by holding the mouthpiece tightly against the

locknut. If the mouthpiece was only partially opened andnot held against the locknut, it would leak at a very highrate. The same test performed on a stock valve revealedsimilar performance.

The astronaut reported satisfactory operation of the urinetransfer system, except for the time required to transfer aurine sample from the suit to the urine " b a g s . This time in-

crement results from the restriction to the flow of fluidcaused by the small bore of the hypodermic needle used inconjunction with the transfer fittings. A quick disconnectarrangement could be used to overcome this problem. At theend of the mission, 1126 grams or 2.48 pounds of urine wereremoved from the urine collection " b a g s .

At 7: 0:18, approximately 300 cc of water was pumped to the

3.86-pound-capacity condensate bag. Later in the mission,the astronaut transferred some additional condensate to thisbag. He noted, however, that he could get very little fluidinto the bag because of back pressure on the pump. He thendrank all of the wateri he could from the 4.5-pound drinking-water tank, relocated the needle to the drinking-water-tankinlet transfer fitting, and pumped several times into thistank. He then relocated the needle to the 3-86-pound-capacitybag transfer fitting and tried again to pump water into thisbag, but with no success. He noted that the pump seemed com-pletely jammed. He then attempted to relocate the needle tothe drinking-water tank transfer fitting, but the transfer-

fitting needle guide of the 3-86-pound-capacity bag unscrewedfrom its base and remained firmly affixed over the-needle.

Therefore, it was impossible to insert the needle into anyother transfer fitting. He then unsuccessfully attempted totransfer condensate into the 1-pound plastic bags. Theseevents occurred at approximately 27:50:00.

The syringe pump stem is made with serrations on its surfacewhich indicate pump stroke capacity in cubic centimeters. Be-cause of the orientation of the syringe in relation"to theastronaut, a considerable side load was placed on the plungerstem each time it was actuated. This load caused a severe

broaching action between the serrated plunger stem and itsmetal guide. The broaching action generated metal chips andslivers, which migrated past the pump plunger, entered thesystem and clogged the transfer needle, thereby preventingfluid flow. The syringe pump operated normally when thesemetal particles were flushed out after the flight.

V.V71!f



Page 5-19

The malfunction of the 3-86-pound-capacity bag transferfitting was caused by a failure to get the proper seatingof a set screw between the fitting parts. The fitting con-sists of a base and a needle guide which screw together toretain the rubber diaphragm. The needle guide external male

threads mate with female threads in the nylon cover shieldof the transfer needle. When the diaphragm was replacedprior to flight, the set screw was not set firmly enough toprevent relative motion between the fitting base and needleguide. Friction between the needle guide threads and thenylon cover threads would have been great enough to cause thefitting-needle guj.de to stay with the transfer needle whendisengagement was attempted. It is possible that the 3-86-pound-capacity bag evidently got some air into it after the evacu-ation had been performed. The expansion of this air duringascent could have reduced the capacity of the bag in orbit.

The total condensate recovered after the mission was 4.2 pounds,assuming that all of the water in the 4.5-pound-capacity drinking-water tank was condensate. The condensate removed from this tankamounted to 2.4 pounds. In addition, 0.7 pound of condensatewas removed from the 3.86-pound-capacity bag, and 1.1 poundsof condensate was removed from the 4-pound-capacity condensatetank.

5-2.4 Thermal and water balance analyses.-

5.2.4.1 Metabolic analysis: The average metabolic heat productionduring the mission may be calculated quite accurately from theamount of carbon dioxide (COp) absorbed in the lithium hydrox-

ide (LiOH) canister. The rate of COp absorption during the

mission, including that experienced prior to both lift-offattempts, was 26l cc/min at standard temperature andpressure (STP). One correction to this figure would be theCOp lost through cabin leakage. With the faceplate closed,

the partial pressure of C0p in the cabin would be zero. How-

ever, the astronaut had his faceplate open at times duringthe flight, and some COp will have been lost directly to the

cabin. It is not possible to determine accurately the averageCOp partial pressure in the cabin during the flight, but the

amount was most likely well below 10 mm Hg. Therefore, theleakage rate of C0p probably did not exceed 2 cc/min (STP).

Assuming a normal respiratory quotient (RQ) of 0.83, theaverage metabolic heat production was 92 kilocalories/hr(kcal/hr), or 370 B.t.u./hr. The astronaut's surface area is

GQNF IDENTIAl r





Page 5-21

calculation of the metabolic rate at specific times duringthe flight was not possible because of insufficient instru-mentation. It is likely that the two readings of 100.0° Fare related to the combined effects of a mild sustained heatstress and moderate physical activity during the early part of

the flight. Further analysis of the thermal condition of theastronaut is not possible because of the absence of detaileddata for cabin environmental conditions and metabolic, rates.It can be said that the deep body temperature did not vary fromnormal sufficiently to impair the astronaut's inflightper-formance. However, the astronaut's skin temperature and sweatrate were at somewhat higher levels than those considered tobe ideal for a suited pilot over this period of time.

The estimate of 11.0 pounds for body-water vapor loss, cal-culated from the thermal balance, can be verified by a calcu-lation of the vapor loss from body-mass balance if all of the

relevant data are accurately known. The major unknown is theintake of drinking water. The astronaut was not certain thathe emptied his ^^-pound-capacity drinking-water tank beforeadding condensate to this tank. The mass balance of theastronaut is given in table 5.2.4.1-1. The maximum possiblebody-water vapor loss is seen to be 10.1 pounds. The actualloss would be less than this value by the difference between4.5 pounds and the amount that was actually drunk from thedrinking-water tank. At recovery of the spacecraft, 2.4 poundsof water were present in the tank, and some of this was condensate.The minimum value for consumption of drinking water from thetank is therefore 2.1 pounds, and the actual consumption was

probably nearer to 4.5 pounds.

The amount for body water vapor loss, calculated from the

mass balance, is less by at least 0-9 pound than the valuecalculated from heat exchange. The heat-exchange value of11.0 pounds is calculated on the assumption that all waterevaporated from the skin came from sweat. During the timethat the condensate trap was turned off,primarily the entireperiod after 12:14:00, free water must have passed into thesuit and been absorbed by the astronaut's undergarment. Thiswater was subsequently evaporated into the ventilating gasand,therefore, reduced the need for sweat production. This

regenerative cooling has been observed during various simu-lations and experiments in the altitude chamber, where thefree water condensed out by the suit-circuit heat exchangerwas allowed to pass back into the suit. Since some watermust have passed into the suit, the total body water vaporloss (maximum, 10.1 Ib) will have been less than the amountpredicted from the thermal balance (ll Ib.). The astronautwas not able to differentiate between the degrees of wetness

INriDENTLIL"



Page 5-22

of his skin and undergarment when the condensate trap wasclamped off and when it was in operation. The pilot did statethat he was somewhat cooler and more comfortable when the conden-sate trap was operating. His undergarment was apparently wet atall times during the flight.

Total water input into the suit circuit also included approxi-mately 1 pound of water liberated from the lithium hydroxidecanister. Total water recovery from the system should havebeen 11.1 pounds.

Of this quantity, the following amounts were recovered andmeasured:

Container

3. 86 -pound -capacity condensate bagsU-pound-capacity condensate tank

Lithium hydroxide canister

Water, Ib

0.71.1

0.2

This tabulation leaves 9-1 pounds of water which can only beaccounted for as follows:

Source Water, Ib

Condensate pumped into drinking-water tank, maximum

Residual water left in the under-

garment and space suit (estimated

minimum figure from previous labora-

tory experiment)3

Water lost into spacecraft from the

condensate tank and leaking valves

Water loss from cabin ECS circuit

during ventilation with ambient

air after snorkels open during

reentry

Residual water in suit environmental

control system

2.0

(b)

(c)

(d)

Not measured.

Values not yet available.

d

'Negligible and cannot be calculated from present data.

Cannot be meaningfully calculated because of inaccuracies.



Page 5-2J

It is emphasized that the computations of thermal and vaterbalance is subject to considerable inaccuracy because ofthe lack of specific measurements of some of the variables.The various locations of water and rates of vater loss givenabove could account for the total water input of a maximum

of 11.1 pounds to the suit circuit of the environmentalcontrol system.

5.2.4.2 Food and water consumption:

The astronaut's total food consumption during the flight wasas follows:

Item

One baconsquare

One containerof dessert

cubes

One -thirdpacket ofdehydrated beefpot roast

Two fruit cakes

One containerof peanutbuttersandwiches

Totals

Time wheneaten, hr:min:sec

06:32:15

0 :53:58

06:32:1526:1 : 2

11: 16: 04

26:14:4228:29:41

28:29:41

Amount,gm

5

88

1

2k

_22_

146

Value,kcal

19

429

32

92

124

696

Although the astronaut was not hungry, he found that the

food he ate tasted good. He wanted to try the freeze-dehydrated foods and juices, but he was unable to put

sufficient water into the food reconstitution bags becauseof leakage from the drinking water valve. The valve wastested after the flight and found to leak freely when thenozzle was at intermediate positions between off and fullyopen. There was, therefore, an inadequate pressure headto fill the food bags with water.







Page 5-26

TABLE 5.2A.1-1.- BODY-MASS BALANCE SUMMARY

Source Weight, Ib

Available for output as vapor

Maximum water intake from drinking

water tank

Survival-kit water

Food

Oxygen (assuming an RQ of 0.8j)

Body weight loss

Total

^•5

1.2

0.3

2.2

7-015-2

Not available for output as vapor

Urine

Carbon dioxide

Total

Difference, maxim-urn water vapor loss

from body

2.6

2.5

5 - 1

10.1









5. .5 Spacecraft window. - The astronaut reported that the spacecraftwindow was coated with some type of substance that caused the

window to have a frosted appearance under oblique lighting con-ditions. This coating was on the inside surface of the outerpane. A postflight analysis of the surface failed to show any-

thing that could cause this coating and the investigation iscontinuing.

5-^.6 Landing-shock attenuation system.- The MA-9 landing-shock attenu-ation system differed from the MA-8 configuration in that thepressure tank of the landing-bag actuation system was altered inshape and moved to the opposite side of the spacecraft to make room

for the hydrogen' peroxide fuel tank that was added to the RCS, andthe check valve on the ground checkout port was, deleted.

5-^.6.1 landing bag.- The landing bag was deployed at a pressure altitudeof 9,500 feet, and the system performed normally, as evidenced

by the astronaut's statements and from postflight examination.The postflight examination of the bag revealed some small tears

and rips, but they were of a minor nature. The straps and cableswere not damaged beyond that normally expected.

5-^.6.2 Ablation shield and main pressure bulkhead. - The ablation shieldappeared to be intact, and only minor circumferential cracks were

noted. The fiber glass protective shield had been scarred by theheat-shield lugs, indicating-that minor recontact, less than inprevious flights, haS occurred. The main pressure bulkhead didnot exhibit any visible damage. A more detailed discussion ofheat-shield performance is contained in section 5-7-

5.^.7 Flotation. - Reports and photographs from the recovery forcesand the astronaut indicate that the spacecraft righted itselfquickly and floated at the proper attitude once the parachuteshad been jettisoned.

5-5 Electrical and Sequential Systems

5.5.1 Electrical system.- The major modifications to the electricalsystem of the MA-9 spacecraft since the previous mission include

the replacement of the two 1,500 watt-hour standby batteries

with two 3 000 watt-hour batteries to increase the availableelectrical energy, the -replacement of the main inverters with

more efficient units which also had improved starting and coolingcharacteristics, the addition of a standby-inverter automatic tonegenerator to indicate automatic switching of the inverter to eithera-c bus, the addition of a switch to allow repowering of certainbuses after landing to permit postlanding recordings of blood-pressure and EGG readings, and the addition of an on-off switchin the flashing recovery light circuit.





5 32 C O N F I D E N T I A L *

circuit. The large inductance associated with the outputimpedance of the inverter, along with the high rate ofcurrent change, resulted in large 600-volt spikes thatshorted the parallel diode (peak inverse voltage rating of300 volts) across the voltage sensing relay. With this diode

shorted, the relay stabilized in its deenergized state andcaused the standby inverter automatically to begin powering theshorted bus at 3 volts. The standby inverter warning light came

on at this time. During this test, the standby inverter wouldnot come up to full power because of the short circuit whichwas present at the power plug. Because this circuit failurecould be easily duplicated in the laboratory following theflight, it is believed a short circuit which prevented theirproper operation could have existed in flight and that no

malfunction of the inverters was present at any time duringthe flight. Additional laboratory tests show that the 1^0-volta-c meter reading reported by the astronaut resulted from

operation of the inverter under the unusual remaining load inthe circuit, which consisted of a 100,000-ohm resistor inparallel with a diode other than the one which was shorted.

5 - 5 - 2 Sequential system.- The major modifications to the sequentialsystem of the MA.-9 spacecraft include rewiring of the pilot'sabort circuitry to prevent inadvertent ignition of the tower-jettison rocket motor, rewiring of the retrorocket arm switchto provide the capability of disarming both the main and

isolated retrorocket squibs with automatic bypass at retro-grade signal, rewiring of the."fire-retro" green telelite tomake it dependent on the ignition of all three retrorockets,

and paralleling the no. 1 and no. 2 main parachute armingcircuits to increase reliability in arming the 21,000 ft drogue

parachute barostats.

The sequential system operated satisfactorily throughout the

mission, except for the premature 0.05g signal from theamplifier-calibrator, which is discussed in section 5-1- Thissignal caused the sequential system to remove power from theattitude gyros, the automatic retrosequence circuitry, theautomatic retropackage jettison circuitry, and the retrorockettelelites. Since these functions were latched out,' the pilotwas forced to use the emergency retrofire circuitry and to

jettison the retropackage manually.

Since the antenna canister was not recovered, the failure ofthe balloon deployment mechanism cannot be explained. Onboardtelemetry verifies that the squib firing relay used in deployingthe balloon was properly actuated. A more detailed explanationof this failure can be found in section 5-8



T'* Page 5-33

5.6 Instrumentation System

The instrumentation system monitors certain parameters relating to

specific equipment in the spacecraft, and, in general, the data are either

displayed to the astronaut or transmitted to the ground, or "both. Thedata transmitted are also recorded on the onboard tape recorder along with

additional data vhich are neither displayed nor transmitted.

5.6.1 System description.- The major changes made to the MA-9 instru-

mentation system include deleting the high frequency telemetry

system and the low-level temperature survey; adding switches to

allow the astronaut to select continuous, off, or ground command

for the telemetry transmitter and continuous, off, or program

for the onboard tape recorder; adding a switch to allow the

astronaut to remove power from the R- and Z-calibration relays

in the event the programer failed to perform this function;7

changing the speed of the onboard tape recorder from ITTinches

per second (ips) to T? ips to increase recording time; and adding

a new type programer which was to perform the following functions:

1. One minute of "on" time of the onboard tape recorder

every 10 minutes of elapsed time, with 15 seconds each of Z- and

R-calibrations during the last 30 seconds of the sixth cycle.

2. Six minutes of "on" time of the telemetry transmitter

together with 15 seconds each of R- and Z-calibrations after a

telemetry "on" command from the ground has been received.

3. Six minutes of "on" time of the C-band beacon together

with 15 seconds each of R- and Z-calibrations after a C-band

beacon "on" command from the ground has been received.

U. Six minutes of "on" time of the S-band beacon together

with 15 seconds each of R- and Z-calibrations after an S-band

beacon "on" command from the ground has been received.

In addition, the astronaut observer camera was replaced with a

l6-mm movie camera, which is shown in figure 5-6.1-1. The

l6-mm camera was a magazine loaded unit which was internally

powered by fourteen R-^-01 mercury cells. The camera could be

hand held by the pilot, mounted on the instrument panel to

photograph the pilot, or mounted at the window to photograph

objects outside the spacecraft. Three magazines of film were

provided. Two magazines were each loaded with approximately

120 feet of Eastman Ecktachrome film and they were to be used

for pilot observer, general, and reentry photography. The





1DENTIAL Page 5

when the astronaut was desuited aboard the recovery ship. Apostflight check of the harness disclosed a failed solder jointfor one of the wires to the side impedance pneumograph electrode.

Initial playback of the onboard tape indicated that the tape

recorder failed to operate in the program mode after 12:18:19.Recorder operation in the continuous and vox-record modes wasnormal throughout the mission. A postflight test of the pro-gramer indicated that the "A" section, which controls theprogram mode of recording, was completely inoperative. Visualinspection of the programer revealed that the "A" section wasinoperative because the mechanical timer was jammed. Thisjamming was caused by a misalined gear, which resulted from abroken gear shaft. The misalined gear is shown in figure 5-6-2-1.

At retrofire, the indicated carbon dioxide partial pressuredecreased rapidly from 3-0 mm Hg to zero. Postflight testing

'revealed a J-percent calibration shift which would not accountfor the inflight occurrence. However, methods have been foundfor. duplicating the sudden drop in PCCU, but these findings as

yet have not been correlated with realistic conditions- whichcould exist in the spacecraft. Investigations to determine thecause of this problem are continuing.

Immediately after retrofire, the indicated temperature of the150 v-amp inverter decreased from 120° F to 86° F in 35 seconds.The indicated temperature then increased slowly, followed bya slow decrease just prior to drogue parachute deployment.About 30 seconds after drogue deployment, the indicated tempera-

ture increased rapidly to 120° F. The temperature of the250 v-amp standby inverter also increased momentarily at thistime by about 6° F. The indicated readings are considered tohave been valid, since the sensors involved were found to beintact and operating normally. Although the possibility existsthat free water in the cabin at retrofire could have beendeposited on or near these sensors, thereby reducing the localtemperature suddenly, this thesis cannot be confirmed and furtherinvestigations are presently in progress.

At 3^-:09:24, there was dropout in the voltage control oscillator(VCO) outputs for the balloon experiment and for the roll,pitch, and yaw stick positions. This dropout lasted until3 :12:51, at which time the stick position outputs returned tonormal. Postflight testing on the stick position VCO's revealed,however, that they were functioning normally. The balloon ex-periment VCO could not be tested because it was in the antennacanister, which was not recovered at landing. Because all of theassociated spacecraft components, except the retropackage andthe antenna canister, which were not recovered, functionednormally during postflight tests, no explanation can be foundfor the dropout of these signals.

• C O N F I D E N T I A L





INTIAL ' Page 5 - 37

The measured weight loss of the heat shield was 15-3^ pounds,

which is comparable to that which was measured following

previous orbital missions.

5-7-2 Afterbody. - No temperature measurements were made on the

conical and cylindrical sections of the spacecraft. Postf lightexamination of the shingles revealed no areas where adverse

heating occurred.

5-7-3 Paint patch evaluation.- Three individual types of white

paint coatings were applied to a Rene ' 4l shingle on the

conical section, as shown in figure 3 • 1-1 • Each patch was

6 inches square and approximately h mils in thickness. The

three types of pigments were titanium dioxide, zinc oxide,

and zirconium oxide. The coatings were applied with a thermal-

drying binder and cured with a special application to obtain

better adhesion and color characteristics . Previous attempts

to obtain similar results were not successful, probably becausean air-drying binder was used. Postflight inspection revealed

that all three coatings were discolored. The titanium dioxide

and zinc oxide coatings remained bonded to the shingle; however,

the zirconium oxide coating exhibited extensive peeling.

(See fig. 5-7-3-1- ) The paint patches did not experience

erosion, as was evident on previous flights. Further laboratory

investigations are necessary before a conclusion can be reached

regarding discoloration and flaking.




Page 5 ~

5.8 Scientific Experiments

A number of experiments and scientific observations were plannedfor the MA-9mission. These experiments were concerned with atmosphericdrag measurements, visual acquisition and tracking effectiveness, various

photographic exercises, cosmic radiation measurements, and general pilotobservations. The results of this experimental effort, which utilizedequipment or materials not normally associated with the basic spacecraftoperation, are briefly summarized in the following paragraphs.

5-8.1 Dim-light photography. • - The dim-light photography experimentwas intended to provide photographic data on two phenomena,which are best observed from outside the earth's atmosphere.These phenomena are the so-called zodiacal light and thenight airglow. Photographs of the zodiacal light would helpto determine its exact origin, geometric distribution, andrelationship to solar radiation and flare activity. Data on

the airglow would provide further information on the solar-energy conversion processes occurring in the upper atmosphere.

Observations of the zodiacal light are partially obscured or

become impossible when viewed through the atmosphere because

of absorption and scattering. For example, measurements of

the brightness of the zodiacal light within 30° °f ^he sun

have not been possible because of the interference of twi-

light.

Space flights provide an excellent opportunity to view the

bright night airglow band in profile. Some data were col-

lected during the MA.-7 mission on the altitude and bright-ness of the airglow layer above the earth by observing and

timing the passage of a star through it. The photographs

planned for the MA.-9 mission were intended to provide\ further

information on the brightness, structure, and altitude of this

airglow band.

The camera used in this experiemnt was a semi-automatic

35-mm hand-held instrument fitted with a fast lens which is

shown in figure 5-8.1-1. The square aperture of the shutter

limited the speed of the lens to an equivalent of f/0.95- Its

controls were simplified so that operation by the pilot in a

pressure suit would be facilitated. Exposures were manually

timed by depressing a trip button for the duration of the

exposure. For aiming purposes, three small supports or

"feet," capped with silicone rubber, were provided to assist

in maintining the camera in a fixed position relative to the

spacecraft window. The camera field of view was 35° total angle,

with little vignetting over the field. Ansco H 529 emulsion

film, having a film speed of 200, was used for the experiment.



Page 5-39

With the spacecraft in retroattitude, the camera was to be

held against the window so that the nominal position of the

earth's horizon would "be 5° below the center of the frame.

Dim-light photography was performed on the night side of the

l6th orbital pass with all cabin lights out. Just prior tosunset,, the spacecraft was oriented into the plane of the

ecliptic and flown on automatic attitude control throughout

the zodical- light photography period. This sequence was

started just after sunset and several photographs, varying

in exposure duration from one to 30 seconds, were taken.

The initiation and conclusion of all exposures were planned

to be voice-marked on the onboard tape to time the duration

of each exposure. A total of 17 exposures were made of the

zodiacal light during this mission.

Upon completion of the zodiacal-light photographs, the gyro

switch was placed in the slave position and the spacecraftwas slowly reoriented to orbit attitude. Throughout the

remainder of the night period, five series, each containing

a 2-minute, a JO-second, and a 10-second exposure, were made

of the horizon and airglow layer. A total of 15 exposures

were made in this series. Two 30-second and two 1-second

exposures were made of the western horizon when the space-

craft was passing over the terminator at sunrise.

Since much of the analysis depends on conducting microdensitom-

etry of the film, it is premature at this time to present con-

clusions for this experiment. A general summary of the photo-

graphs obtained, however, is available. In the zodiacal-lightsequence, the first two frames showed exposures of dim light,

and all exposures contain stars. The photographs, however,

are generally underexposed and will be of limited value for

analysis. Because of the strong gradient in the zodiacal-

light intensity near the sun, either a small delay in begin-

ning the photography sequence or improper camera aiming would

result in underexposure of the film. A delay of 2 minutes in

beginning the exposures, an 8° e'rror in aiming the camera, or

a combination of the two could decrease the light intensity by

a factor of 10. Because of the known variance in spacecraft

attitude when in the automatic control mode, some attitude

misalinement would naturally be expected. The astronaut also

stated that the message he transmitted over Zanzibar on the

16th orbital pass delayed him somewhat in preparing for the

dim-light experiment, and preliminary data indicate that the

sequence was not begun until 1 minute 0 seconds after sunset.

However, the star patterns appearing in the zodiacal-light

photographs should provide a reference from which camera

sighting can be accurately reconstructed.



P a g e 5 - 4o C O N F I D E N T I A L ^>

The exposures in each of the five series of -airglow photographs

were all of usable quality. Figure 5-8.1-2 is a representative

photograph of this series, shoving the thin "band of the airglow

layer just above the horizon. Some of the 2-minute exposures

are slightly blurred by rates in roll and pitch, but most of the

10-second exposures are quite sharp, vith the airglow layerlocated near the base of the photograph. Preliminary measure-

ments of the angular thickness of the airglow layer indicate

that the layer is about 5-5°in thickness. The four exposures

made at sunrise were all overexposed.

5.8.2 Ground-light experiment.- The ground-light experiment was

designed to evaluate the operational problems associated with

visual acquisition from space of earth-based lights and to

compare the observed brightness with predicted values. This

type of information is necessary to determine the feasibility

of using ground lights as navigation fixes during future space

missions.

The light assembly used for this experiment was a pulsed xenon-

arc type consisting of three sections of six lamps each as

shown in figure 5-8.2-1. The lamps were mounted in a shallow

open-top box above a polished reflective surface. The circuit

was operated by a 50-cycle three-phase electrical power supply,

and the lamps in each section flashed 50 times a second with

an input of 27-5 watt-seconds per flash per lamp. Based on a

if-0-kilowatt nominal input to the lamp power supply, the com-

puted average intensity of the light was estimated to be

approximately 120,000 candles in the hemisphere above the light.

The light was located in the Republic of South Africa, 6 miles

east of the town Bloemfontein, which is at 29°06' S latitude,

26cl8' E longitude, and k,k^>k feet in elevation. With the

spacecraft at an attitude of - 0° in pitch and 0° in roll and

yaw, the light would first come into view at about 8:23:00 and

a slant range of 320 nautical miles with an estimated inten-

sity of about a 2.8-magnitude star.

The lamp was turned on at 8:21:56 and turned off at 8:25:06.

The weather was reported on the ground as having been completely

clear at that time. The light was seen by the astronaut prior

to moonrise during his pass over this area. By moving away from

and back to a fixed attitude, the pilot was readily able to

reacquire the light. The pilot estimated that when the light

came into view, it was equal to about a third-magnitude star.

It was in sight for 30 to ^0 seconds before it faded out at

which time the intensity of the light was between Uth and 5th

magnitude in brightness. The experiment produced approximately

the same sighting results as predicted.

CONFIDENTS



• J C O M l D C P f T I A L ^ge 5 - M

The astronaut considered the light of sufficient brightness

. to be used as a navigation landmark if adequate sighting

information were to be made available. He stated, however,

that the distinctive U-shaped light pattern of the town of

Bloemfontein was a very helpful factor in identifying the

light; the light from the town was about as visible as thetest light and both faded from view at the same time. He

stated that a flashing light would have been much more dis-

tinctive, and he "believed that a pattern of flashing lights

would have been even more distinguishable.

The pilot's observations of the ground light also show that,

for any use of a lighted landmark as a navigation aid, the

rapid passage over the ground (involving angular tracking

rates of as high as l80°/minute) requires that instrument

readings be made very conveniently and rapidly. An attempt

to use the extinction photometer for this experiment was un-

successful.

5.8.3 Radiation studies.- Instrumentation was placed aboard the

spacecraft to measure and record the amount of radiation

encountered on the flight. These measurements are needed to

verify theoretical flux estimates and shielding calculations.

and to ascertain the dose received by the astronaut. Two

major types of radiation are present within the Mercury

orbital altitude range: Protons which are naturally present

in the lower Van Allen belt and electrons which were produced

by decaying fission nuclei from the atomic explosion of July

1962. These radiations are most pronounced for the space-

craft's orbit in an area over the South Atlantic between thecoasts of South America and Africa because of an anomaly in

the earth's magnetic field at that location.

Two Geiger counters were mounted on the retropackage (see

fig. 5-8.3-1) to measure the electron flux incident on the

vehicle. One counter surveyed a hemispherical area about an

axis alined along the 195° yaw and -lit-00pitch attitudes,

relative to the spacecraft axes. The other counter was

collimated to view a solid angle of approximately 0.8 steradian

along the spacecraft's Z-axis. The latter counter, coupled

with vehicle attitude information, was included to provide

Information on the directionality of the electron flux. Thetubes in both counters had a mass concentration in the wall

i~\

of 90 ilO mg/cm and were filled with a low-pressure, neon-argon,

and halogen mixture. The uncollimated tube had, in addition,

a 1-mm tungsten shield to allow passage of primarily high-energy

electrons. Although Geiger tubes are sensitive to protons, the



P a g e 5 - te C O N F I D E N T I A L

count rates recorded were expected to result primarily inelectrons, since the electron flux is several orders of magni-tude higher than proton flux in the geographical region inquestion. A high-voltage power supply maintained an operatingplateau of approximately 550 volts. Power to the tubes was

controlled by a toggle switch on the main instrument panel.The voltage output, proportional to the count rate, was multi-plexed and stored on the onboard magnetic tape.

Radiation inside the spacecraft was monitored by a self-indicating ionization chamber placed on the interior of theegress hatch, which represents a region of minimum shielding.The ion chamber, combined with outside flux measurements,was intended to yield information on the minimum shieldingcharacteristics of the spacecraft and permit an estimate ofthe dosage received by the astronaut. Electron-sensitive filmbadges were placed in the astronaut's helmet and on his under-

wear, in the vicinity of the chest and the right thigh, tomeasure the received dosage. A nuclear emulsion package wasmounted behind the instrument console to register the protondosage.

Since the data were recorded on the onboard tape, it was neces-sary to have the tape recorder on during periods of measurement.During passage through the magnetic anomaly for orbitalpasses 5, 7, and 8, the switch was turned to the "on" position,and the magnetic tape recorder turned to "continuous." Back-ground readings were obtained on the l6th, l8th, and 19th

orbital passes.

The ion chamber was placed on the hatch within 1 hour afterlift-off and remained there until it was stowed just priorto retrofire. A postflight reading of this sensor was obtainedon the ground.

For the other radiation monitors, no astronaut participationwas required. These items were recovered and are being analyzedby the U. S. Naval School of Aviation Medicine for evaluation.

Appreciable count rates from the Geiger tubes were obtained,as expected, on the 7th orbital pass. The spacecraft was in

ASCS orbit attitude during the pass through the anomaly andshould give information on the electron flux and directionsrelative to the spacecraft. However, at the time of thiswriting, considerable data reduction and interpretation remainto be accomplished before any quantitative statements can bemade on the electrons in this anomaly.





Page 5:- kk C O N F I D E N T I A L - *

The infrared spectrum from 660 to 900 millimicrons wavelength:was divided into three parts by inserting filters in the camera'so that sections of the image were photographed through each of-,

the three filters. Kodak Wratten gelatin filters measuring2-1/4 by 3/4 inches each were mounted in vertical strips in a '

thin metal holder, which was located in front of the film plane.The filters are listed in the following table. The lower boundaryof each spectral region is defined by the filter transmission andthe upper boundary is limited by the film sensitivity.

Wavelength,

660

730

810

millimicrons

to

to

to

900

900

900

Filter

W-70 and 1

W-88A

W-87C

and

.0 neutral density

0.9 neutral density

A neutral density filter was added to the W-70 and W-88A filtersections to obtain the same light attenuation through thesefilters as was obtained through the more opaque W-87C filter.The filters appear in the photographs in the order listed inthe table, with the W-70 filter section on the left, theW-88A in the center, and the W-87C on the right.

Kodak high-speed infrared film was used and coated on a regular

base support (0.005 inch thick). The film is sensitive throughthe visible region of the spectrum and in the infrared to approxi-mately 900 millimicrons, with the highest sensitivity in theregion from 770 to 84o millimicrons. The filters absorbed allthe light from the visible spectrum and transmitted to the filmonly the infrared wavelengths. The ASA daylight-exposure indexis 80 without filters. A shutter speed of 1/125-second and alens aperture setting of f/5.6 was used for these exposures.

A series of l6 infrared photographs were taken by the astro-naut over the southern part of the United States and the southernpart of Africa near the end of the 17th and beginning of the

l8th orbital passes, starting at an elapsed time of 26:38:00.Figure 5-8.4-2 is representative of the photographs obtainedfor this experiment. Geographic locations can be determinedaccurately from landmarks in the first ten pictures where cloudsdo not obscure the earth. These pictures showed that cloudscovered most of the Pacific coast of Worth America. Cloudiness,which was associated with a low-pressure area over Oklahoma,covered most of the lower Mississippi River Valley region and





P a g e 5 - k6 " CQNPIDEJmAL

space missions. Results from the experiment conducted during the

MA-9 flight were intended to extend the data previously obtained

during the MA-7 mission on the relative elevations of the earth's

limb as -photographed through blue and red filters. Past results

indicate that this variation in earth-limb height was dependent

upon the scattering angle of the incident sunlight. Therefore,a time-extended series of pictures embracing most of the daylight

period of an orbital pass was desired. It was also desired to

obtain a group of photographs taken over a brief time.period

in four quadrantal directions relative to that of the sun. Ac-

curately timed photographs of the earth horizon including the

setting moon were also requested to provide a reference for

measuring the altitude or elevation of the limb.

These photographs were made with the Hasselblad camera, pre-

viously described in section 5.8A. A lens setting of f/8 and

a shutter speed of 1/125 second was used. A filter sandwich

consisting of a no. 92 (red) Wratten filter central panel andtwo no. i)-7B(blue) filter side panels were mounted immediately

ahead of the film plane. A narrow, opaque bar extended through

the center of the red filter to hold the filter flat. Linagraph

Shellburst film was used for the experiment.

A planned series of photographs of the earth's horizon through-

out the daylight portion of an orbital pass was not performed

during the 21st orbital pass, as programed, because of the space-

craft control system malfunction. The quadrant photographs and

photographs of the moon-earth limb were taken as planned. A

total of 11 exposures were made, 8 quadrantal photographs (2 of

each quadrant) and 3 moon-set earth-limb photographs.

Because of an unforeseen difficulty in loading the modified

camera back and a consequent inadvertent exposure of the film

to light from a red safe-light, some background "light-struck"

damage was incurred by the film prior to flight.

Careful initial inspection of the film indicated that six of

the eight quadrantal exposures and two of the three moon-set

pictures are usable for quantitative study. The film includes

two good sets of step-wedge sensitometric exposures for cali-

bration, one before the flight and one after. Correction for

the inadvertent background can be made, at least where it isof moderate density.

Since these densitometric exposures were made with a calibrated

source, a radiance value for the observed limb can be eventually

obtained from the data. The MA-9 results in this regard, con-

stitute a significant advance beyond the photography of the

MA-T flight.









P a g e 5 - 50 GOMlDEMnAL

During the third night period, at 3 hours 25 minutes after

deployment and at a range calculated to "be betveen 9^ and1 ^

11 nautical miles, the light was rated as very veak and just

"barely discernible. This rating indicated again that thebeacon was not as bright as expected. The astronaut alsoestimated distances as well as brightness; and since therewere no distance cues other than those related to previousaircraft sightings, he consistently estimated distances thatwere too great.

The sightings were made with all spacecraft internal lightsturned off. An attempt to sight the beacon during the day-light period after the first night period, when the calculatedminimum range was less than 2 nautical miles, was unsuccessful.Although, at other times, stars were seen during the daytime,

none were seen during the daylight period in which the astro-naut was trying to acquire the light.

In general, it was found that the flash of the light made iteasily distinguishable from stars. The "beacon's intensityappeared to be adequate for acquisition up to distances ofabout 8 nautical miles at night.

5.8.7 Tethered balloon experiment.- A JO-inch-diameter mylar spherewas packaged in the antenna canister of the spacecraft and wasto be ejected, inflated, and tethered at the end of a 100-footnylon line for one orbital pass. However, the balloon failedto deploy and the experiment was not accomplished. This experi-ment was intended to provide further information about theatmospheric density through the altitude profile of the Mercuryspacecraft, which ranged from 8k to 1^6 nautical miles. Atmos-pheric density was to be computed by determining the drag ofthe balloon, as measured by the pull of the tether on a strain-gage balance located at the bottom of the balloon canister. Theequipment configuration for the experiment was nearly identicalto that included in the MA-7 spacecraft for a similar experimentexcept that the balloon was painted fluorescent orange and it

was constructed by laminating a T--mil-thick metalized plastic

sheet to a -mil-thick clear plastic -sheet. The l6-mm camera

(see fig. 5-6.1-1) was mounted in the spacecraft so that photo-

graphic coverage of the balloon's motions could be obtained.





Page 5-52

tend to fall in patterns that would "be easily discernible to

the pilot. Furthermore, since the spacecraft was drifting, theamount of light falling on the window should be varying con-tinuously, but the pilot stated that the lightness of the skyappeared uniform. Therefore, it appears unlikely that thisobservation can be accounted for by scattered light on -th e

window.

The best explanation for this brightening of the sky is theo

presence of a general 6300 A emission region above the space-o

craft. The height of the 6300 A dayglow is not know, but mayoccur in the F-region of the atmosphere which is at an altitudeof about 250 kilometers. Preliminary estimates indicate that

o

the visible emission of the 6300 A dayglow could increase thebrightness of the day sky about 3 times above the level of the

night sky. This hypothesis agrees reasonable well with the

pilot's estimate of the reduction in light intensity of starsvisible during daylight.

All of the astronauts have reported the presence of a so-calledhaze layer on the night side of an orbital pass. This termappears to be a misnomer, since the phenomenon probably involvesa region of concentrated airglow emission in the 100-kilometerregion of the atmosphere. The effect of weakening a star'sintensity as it passes through this region could be caused bya loss of visibility, with respect to the increased brightnessof the airglow foreground, as readily as if it were caused by

the presence of a haze or attenuating layer. Present evidence

supports the former explanation, since measurements of the heightof the layer agree with that of the night airglow layer.

The pilot also reported observing a brownish glow in an east-northeasterly direction above his orbital altitude at an elapsedtime between 17:25:00 and 17:30:00. At this time, the space-craft was just above the east coast of South America in thevicinity of Belem, Brazil. Thus, the phenomenon occurred inthe tropics not far from the geomagnetic equator. The glowappeared as a patch with an approximate angular size equal tothe spacecraft window, rather than a layer, and the observationseems to be similar to that reported over the Indian Ocean during

the MA-8 mission. Postflight calculations indicate that theo

phenomenon was caused by the integrated effect of a 6300 A

cloud of about 3,000 kilometers in horizontal extent. If thetypical vertical thickness is assumed to be about 50 kilometers,then the integrated intensity of the cloud seen edge-on would beabout 18,000 rayleighs, about the intensity required to give avisual impression of a light brownish nature.



Page 5-53

The pilot also reported that at about 20 seconds after sun-

set, a whitish arch extending some 15° or so above the horizon

appears above the sun. Two possible explanations of this arch

exist:

1. It could be the outer corona of the sun.

2. It might be associated with a concentration of dust

around the earth - a sort of terrestrial "zodiacal light."

Approximately 1 minute after sunset, the pilot detected the

zodiacal light. At this time, the spacecraft window was pro-

bably significantly illuminated by the bright remnants of the

orbital twilight. The ability of the pilot to detect the

zodiacal light above the window foreground suggests that it

must have been brighter than the orbital twilight. This is

as it should be at angular distances from the sun of 6° to

10°, which was probably the minimum observed distance. Thefact that the light was concentrated along the ecliptic

established it as truly "zodiacal light."

Thirty general purpose color photographs were taken on this

mission by using the 70-mm Hasselblad camera described in

section 5-8.4. The exposures were made without a filter on

ultraspeed Anscochrome (FPC 289) by using a lens setting of

f/l6 and a shutter speed of 1/250 second. Photographic cover-

age of portions of northern Africa, Arabia and the India-Burma-

China area was obtained. A number of very good exposures of the

Himalayas with excellent resolution were obtained. Figure 5-8.8-1

is an example of these photographs. Some photographs of thewestern Pacific and of the Indian Ocean areas were also taken.

In general, the quality and resolution of these photographs

were excellent, and the resolution of terrain features of

the best of the MA-9 photographs compare favorably to that

of the Viking infrared photographs. Interestingly enough,

however, no cultural features whatsoever can be seen on these

color photographs. Even the city of Calcutta, India, could not

be identified, although the quality of the photograph was ex-

cellent. A complete analysis of these photographs will require

a considerable amount of time; however, a preliminary analysis

indicates that this film-camera combination is capable of

yielding photographic information which could be useful for

geological, topographical, and meteorological purposes. All of

the photographs, even if slightly overexposed, contain some usable

information.

C O N F I D E N T I A L -





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6 - 2 C O N F I D E N T I A L

probe in the sustainer-engine lubricant tank, vhich was also installed in

the engine compartment, was used for the first time in this location. The

probe sensed temperatures of 50° F at lift-off, 70° F at BECO, and 9 ° F

at SECO. This transducer was shielded from thermal radiation and provided

a temperature study of the lubricant as it is consumed.

6.3 Propellant Tanking

The tanking of EP-1 fuel to the 100-percent probe resulted in a total

fuel quantity at 80° F of 11, 28 gallons, which corresponded to an indi-

cated weight increase of 76,500 pounds. Following the countdown for lox

tanking, a drop in fuel temperature and a resultant density increase re-

quired the addition of 350 pounds of fuel to reestablish the 100-percent

level. Final fuel weight, adjusted to include the weighing system cali-

bration, gravity correction, and pressurization gas weight, was 7 ,900

pounds.

Lox tanking to the 100-percent probe required 17 ,000 pounds oflox.

Subcooled topping was added to maintain an indicated lox weight constant

at 17^-, 100 pounds. After securing lox tanking, 500 pounds of lox were

calculated to have been lost through the vernier-engine bleed valves and

the start tank vent.

Following final pressurization, the lox level was above the high-level

probe for approximately 10 seconds. However, this level was between the

high- and low-level probes at ignition because of boil-off. The total

weight of the spacecraft and launch vehicle at ignition was 326, 00 pounds.

6.k Propellant Utilization

The propellant utilization (PU) valve operated normally within the

dynamic range up to approximately 35 seconds prior to SECO. At this time,

the valve followed corrections to near-nominal values over a 10-second

period prior to SECO. At this point, the PU valve followed the bias in-

troduced by the capacitancepad.

The fuel and lox-head pressure-sensing ports were uncovered at 3-25

and 6.85 seconds, respectively, prior to SECO. Propellant residuals at

SECO were calculated to be 83.6 pounds of lox and 5 0.6 pounds of fuel.

6.5 Pneumatics

The lox-and fuel-tank ullage pressures were within normal operating

limits throughout the flight. Booster-tank helium-bottle decay was normal

and the sustainer-tank helium bottle maintained adequate pressure.






p a g e 6 - k G O f f r i D C H T I A L . » .

6.8 Guidance

The launch-vehicle radio-guidance system performed veil and guided

the sustainer stage to near-nominal cutoff conditions that were veil

within the acceptable limits. The radar data were of good quality and

exhibited lov noise. The errors at insertion were 1.3 ft/sec low in velocity,1,333 feet high in altitude, and 0.0031° high in flight-path angle.

The guidance-system configuration was similar to that used in the

MA-8 mission except that the duration of the second-stage pitch program

was reduced by 5-0 seconds, resulting in 10° less pitch-down maneuver, and

the ground guidance computer equations were changed. The reduction in the

pitch programer duration gives a more efficient trajectory. Hovever, the

MA-8 capability of reaching acceptable orbital insertion conditions for a

near-nominal trajectory with a loss of closed-loop launch-vehicle guidance

steering vas not maintained.

A somewhat depressed trajectory was flown throughout the flight.The BECO discrete signal was 0.5 second early, the velocity was high, and

the altitude was low at staging. These conditions resulted in SECO being

1. U seconds early. The radar elevation angle at SECO was a nominal1. °.

The guidance system acquired the track beacon of the launch vehicle

in the first radar cube, and lock was continuous from 00:01:04.0 to

00:06:07.1 (6k sec after SECO). Rate lock was continuous in all functions

from 00:00:58.3 to 00:05:55.7 (53 sec after SECO), except for 0.1 second

of bad central rate data 3-7 seconds after BECO. These bad data were caused

by the jettisoned booster stage assembly's passing between the ground radar

and the airborne transponder. The bad data had no effect on the flight of

the launch vehicle, since guidance steering was not initiated until

22.8 seconds later.

Closed-loop radar guidance steering started at 00:02:38.8 with a

25-percent positive pitch rate and a 30-percent positive yaw rate. This

yaw right maneuver corrected a booster-stage roll program error. There-

after, the commands were smooth and small until 5-5 seconds prior to SECO

when 25-percent pitch steering commands were sent to the launch vehicle

to correct the error in flight-path angle. Noise in the steering commands

was lower than for all previous Mercury-Atlas orbital missions.

The ground guidance computer equations had been changed since MA-8

to extrapolate radar data better and to smooth out radar noise. Guidance

initiation was started on slant range rather than on time after BECO.

Other than these changes, the application of these equations was the same

in the normal data mode. The reason for these changes was to give better

steering commands in the event of noise resulting in the loss of radar

data.

In figures 6.8-1 and 6.8-2, the velocity and flight-path angle are

shown in the region of sustainer engine cutoff. The launch-vehicle data



Page 6 - 5

are shown in figure 6. 8-1, and the range safety impact predictor computer

(I. P. 709*0 data are shown in figure 6.8-2 to illustrate the data quality

during the time of the go no-go computations. Both data sources exhibited

low noise and the average of these values agrees with the actual flight

conditions. One launch vehicle guidance data point was not received by

the Goddard computer at 00:05:08.07; however, a 20-point average was stillused. Maximum peak-to-peak deviations in the launch-vehicle guidance data

were about one-quarter of the magnitude experienced on MA.-4 and MA-8 and

about one-half the noise level of MA-5, MA-6,and MA-7. The IP-709 data

noise level was about half that experienced on MA.-4 and MA-8and about the

same as MA-5, MA-6,andMA-7.

In figure 6.8-3, "the variation of flight-path angle with velocity is

the type of display used by the Flight Dynamics Officer in the Mercury

Control Center for the orbital gono-go decision. Both the launch -vehicle

guidance and HP-709 data indicated acceptable conditions.

6.9 Abort Sensing and Implementation System

The ready status of the abort sensing and implementation system was

properly established at 0.8 second prior to lift-off. Telemetry records

indicate that the pressure switches operated correctly. Wo system

parameters reached the abort level, and no abort command was generated

during powered flight. After SECO, the system indicated an abortcon-

dition, which is an expected occurrence and results from a decay in the

engine fuel pressure. An incomplete listing of the proximity of monitored

parametric values to the specified limits for the MA-9abort system is

presented as follows:



Page 6-6

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values

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7-0 ASTRONAUT ACTIVITIES

This section contains a detailed analysis of the physiological

responses of the astronaut, an evaluation of his performance in com-

pleting the prescribed flight plan, and a flight report "by the astronaut,The aeromedical analysis documents the preflight medical activities,

the inflight responses, and the postflight examinations; and conclusions

derived from the analysis of these data are presented. The astronaut's

performance is evaluated from the standpoint of his ability to maneuver

the spacecraft, manage the operation of the onboard systems, conduct

scientific experiments, and in general complete scheduled .activities.

This section is concluded with a narrative account by the astronaut of

his flight experience and his evaluation of many operational aspects of

the mission.

7-1 Aeromedical Analysis

7.1.1 Introduction.- The MA-9 flight represents another step in

the serial extension of the observation of man as he functions

under the physical stresses of zero gravity, decreased ambient

pressure, and a pure oxygen environment, and under the psychologica

stress of this remarkable experience with the multitude of

risks involved.

The step-wise extension of the mission durations produces

an increasing understanding of man's adaptation to, and

toleration of, this new environment. Application of this

information enhances the confidence in and understanding of

future missions of greater duration. The purpose of this

section of the report is to present the qualitative and

quantitative medical analysis conducted within the Project

Mercury mission framework and thus to indicate the present

state of medical efforts in the area of manned space flight.

The report is presented chronologically in the three categories

of preflight, inflight, and postflight findings. Within each

of these broad areas the data are presented in two parts:

results of clinical studies, both of a conventional nature

and selected special procedures; and results of biosensor

monitoring studies. The concluding portion presents a series

of special medical studies for the MA-9 mission.

The intimate interaction of the pilot and his environmental

control system requires that the discussion of the Life Support

System (Section 5-2) be considered in conjunction with the

Aeromedical Analysis.




Page 7-2

7-1.2 Preflight observation.-

7.1.2.1 Clinical data: Data were evaluated from very thorough medical

studies of the pilot, Astronaut L. Gordon Cooper, Jr., conducted

immediately prior to his selection for astronaut training in

1959 and from annual examinations since that date. Medicalexaminations were also conducted both before and after six

preflight spacecraft checkout tests and a session in the

Mission Control Center procedures trainer, all of which required

the pilot to wear the full-pressure space suit. Special

examinations to assess the pilot's fitness for flight were

conducted 11 and 3 days before launch. The latter examination

conducted on May 12, 1963, designated the "Comprehensive

Medical Evaluation," was conducted by specialists in internal

medicine, ophthalmology, neuropsychiatry, radiology, and

aviation medicine. The NASA Flight Surgeon who had examined

the pilot for most of the preflight activities conducted the

final preflight medical examination on launch morning. Thepreflight aeromedical procedures and examinations are listed

in table 7.1.2.1-1.

The astronaut's pertinent medical history is summarized here

as background medical data. At age 6 he contracted pneumonia

of the left lung, followed by empyema and requiring surgical

drainage. He has residual non-symptomatic adhesions and

thickening of the left lateral pleura and there is a well

healed linear 5 cm scar in the skin at the 9~th posterior

intercostal space at the surgical drainage site. Neither

complications nor sequellae have occurred as a result of this

condition, and repeated pulmonary function studies have beennormal. On November 3? I960, a cholecystectomy and incidental

appendectomy were performed following X-ray demonstration of

cholelithiasis. Full recovery was prompt and no food

intolerance or other sequellae have occurred. He has moderate

seasonal rhinitis relieved by small doses of common

antihistamines. In the three months prior to the MA-9 flight

no antihistaminic medication was required. The pilot

exhibited a gastrointestinal intolerance, evidenced by

cramping abdominal pains, to an oral test dose of morphine

sulfate. This reaction is seen in a small percentage of

normal people. Other than the foregoing items, the pilot's

medical history is not significant.

In addition to examinations by physicians, baseline clinical

evaluations included an audiogram, an electrocardiogram, a

chest X-ray and laboratory studies of blood and urine. The

results of these evaluations are found in tables 7-1-2.1-2

through 7.1.2.1-5. For the 3 months prior to the flight,

the pilot continued in excellent health with significant

abnormalities.




O T I D C M T 1 A L * p a g e 7 3

Close supervision of the pilot's food intake began 7 days before

the planned flight with special preparation of a normal

balanced diet. To reduce the need for defecation during the

mission, a low residue diet was followed for 4 days before the

launch. This diet was well tolerated, although the pilot did

mention that appetite satisfaction was short-lived following

the low residue meals. The constituents, both qualitative and

quantitative, of this diet are presented in table 7.1.2.1-6.

In the month prior to flight, he maintained his physical

fitness by daily distance running and calisthenics.

The aeromedical countdown, table 7-1-2.1-Tj presents the pilot's

activities on launch morning. Compared with the pilots in

previous orbital missions, the MA-9 astronaut spent the

shortest time (5 hours and 1J minutes) from awakening to

lift-off. In spite of a last minute requirement for an earlier

insertion, all elements of the aeromedical countdown were

accomplished satisfactorily without altering the planned

wake-up time.

The final prelaunch examination showed a healthy pilot who was

ready for the mission. Two minor discrepancies were local skin

erythema at the biosensor sites and moderate erythema,

edema, and tenderness of the skin over the right sacral

prominence. He frequently demonstrates a skin reaction around

the sensors for 2k to j6 hours after application, despite the

use of microporous surgical tape for fastening these sensors.

It should be noted that these sensors were in place for

7 hours during the canceled launch on the preceding day. The

skin findings over the sacrum are frequently present following

prolonged periods of 4 or more hours on his back in the couch.

On the night before the postponed launch of May 1 -,19 3, the

pilot slept well for about two hours and then dozed fitfully

for another 3 - hours. He had several dreams related to

problems with the use of the flight plan and the inflight

medications and also later reported he did not feel rested.

However, on the night before the successful launch, he slept

well for 6 hours and no dreams were recalled. Although he

did become sleepy during periods of relative inactivity,such as in the transfer van, he did feel adequately rested

on launch morning. At no time was a drug administered to

induce sleep.

7-1.2.2 Biosensor data: The sources of detailed biosensor data are

outlined in tables 7.1.2.2-1 and 7.1.2.2-2. These sources

include dynamic tests for evaluation of general physical





Page 7-5

The preflight biosensor data are presented in tables 7.1.2.2-1

to 7-1.2.2-k. The analysis methods used were both manual and

automatic.

All respiration minute rates were obtained by 30-second

counts made from the continuous direct-recorded analog signal,with sampling intervals either every 3 or every k minutes.

Heart rates were determined in the same manner by counting

manually; those sets can be readily identified by the

relatively low number of values used. The automatic analysis

utilizes a general purpose computer to determine the intervals

between all the R waves of the EGG complexes in the record,

and the reported values are computed from these determinations.

The automatically reduced rates are readily identified by the

large numbers of values . The validity of both of these methods

has been substantiated by repeated cross-correlation of results

during the two years of development of the analysis program.

Although the data analysis format was arbitrarily selected,the results are fully reproducible and appear to be adequate

for the present medical requirements. All blood pressures

on record were incorporated in the tables.

The most significant aspect of the preflight data is the

rather wide range of mean rates, particularly heart rates,

which limits the establishment of expected or so-called normal

responses . This wide variation is a common phenomenon among

healthy individuals and, in this regard, makes it difficult

to predict accurately the responses of such a person under

conditions similar to those during which the background data

are derived.

The EGG from the preflight observation period was scanned

repeatedly by numerous observers. The collective opinions

were that marked normal sinus arrhythmia was present with

frequent occurrences of a wandering cardiac pacemaker. At

times, sinus node suppression was sufficient to allow

activation by the atrio-ventricular (A-V) node with escape and

fusion beats. This occurrence was identified by both biphasic

and negative P waves of decreased amplitude, and on occasion

by changes in the ventricular complexes. There were numerous

such beats noted during the countdown of the postponed launch,

and there was one brief episode of nodal rhythm during this

period. These data are illustrated in figure 7.1.2.2-3 and

7.1.2.2-^4-. There was sinus bradycardia, which, at times,

was followed by a sinus-generated beat and, at other times,

an A-V nodal-generated escape beat. Other infrequent rhythm

alterations were premature atrial and ventricular beats.

These preflight data were collected in order to establish

the baseline physiological responses of the MA-9 astronaut

specifically using the flight biomedical instrumentation.

O O M F I D E M T I A L



p a g e 7 - 6 C O N F I D E N T I A L

7.1.3 Flight observations.-

7.1.3.1 Biosensor data: Inflight biomedical monitoring spanned a

time interval of 3^ hours, l6 minutes,, and ^3 seconds on this

flight. Continuous onboard recording included the first

1 hour and 35 minutes and the last 10 hours and 5 minutes offlight time until bioplug disconnect. Flight data were

programed to be intermittently recorded for 1 minute of every

10 minutes between 1 hour and 39 minutes elapsed time and

23 hours and 32 minutes elapsed time. Recording of physiolgical

data through the mid-portion of the flight was erratic and

did not follow original plans because of a malfunction of the

tape-recorder programer which occurred at approximately

12:00:00 and continued throughout the flight. Data during

the final portion of the flight, from 24:00:00 until landing,

were made possible because the failed programer was overriden

by the astronaut's selection of continuous recorder operations.

However, sufficient data points were obtained for confidentextrapolation of trends of physiologic values during such periods

by the astronaut's voice contacts with the ground, use of the

vox-record actuation of the tape recorder, or turning the

tape recorder temporarily to continuous to document certain

inflight experiments. During the period when the astronaut

was resting quietly or was asleep, essentially no medical

data were obtained on the onboard recorder; consequently mean

heart rate values for the entire duration of the flight are

probably biased on the high side of a true mean. Data from

the onboard recorder have been supplemented by data obtained

during network station passes throughout the mission, and an

exceptionally valuable short period of recording was obtainedonboard the carrier during egress of the astronaut. The

inflight responses are summarized in tables 7-1-2.2-2 and

7.1.2.2-4. Hear rate response, including mean rates, was

obtained through a computer reduction of the inflight data

from the onboard tape recorder.

Respiration rates were obtained by the manual counting of

30-second periods every 3 minutes during the period of

continuous recording, and from 30-second averages taken at

all other short intervals when data were available from the

onboard tape recorder. Blood pressures were obtained according

to the flight plan with only very minor variations. Thesevalues were with few exceptions not recorded on the onboard

recorder since the astronaut was generally quiet while sending

the blood pressure, and therefore the tape recorder was not

operating. However, the values were received at ground

stations in every instance and read in real time by medical

monitors. The readings were subsequently verified by





p a g e 7 8

When the flashing light was deployed at about 3:26:00, his

heart rate rose to a sharp peak of 13 - teats per minute and

then promptly declined to 95 teats per minute while the

pilot was maneuvering the spacecraft in an attempt to sight the

flashing light. Heart rate remained stable around 80 beats

per minute throughout the remainder of the first 8 hours inspace except during periods when the astronaut announced

on the tape that he was performing some specific exertion

such as emptying the condensate tank or removing equipment

from the equipment 'kit. During these intervals, rates would

increase to values from 100 beats per minute to as high as

130 beats per minute for very short times.

At 8:25:00, the pilot specifically mentioned struggling with

his writing desk. At this time, his heart rate was seen to

rise to 96 beats per minute and then promptly settled back

to its resting rate of about 80. The longer period of

observation and the opportunity which this flight affordedto correlate pilot activities with heart and respiratory rates

permits tentative appraisal of the effect on these rates of

exertion in the orbital spacecraft based on similar physical

exertion under equally cramped circumstances at Ig. There

does not appear to be a significant difference in terms of

pulse rate and respiratory rate response in the two situations.

This impression was further borne out in the two planned

exercise periods where there was consistent similarity

between the response to exercise in orbital flight and the

response to exercise in preflight practice sessions, as

shown in table 7-1-3-1-1.

The respiratory rate sensor became unreliable during the

flight. The failure was subsequently traced to a separation

of the electrode lead wire from the electrode, which was

attached to the left lower chest. The first sign of respiration

sensor failure occurred at 7:08:00, and throughout the

remainder of the flight, the respiratory rate recording was

intermittent. Sometimes it appeared to trace a faithful

replica of the pilot's breathing, but at other times it

was entirely unreliable or without apparent relationship to

respiration. The respiratory rates during the last portion

of the flight are tentative rates based on the appearance

of the pneumograph waveform during periods when evidenceavailable indicates it was following changes in thoracic

volume. Typical signals for properly operating biosensors

are illustrated in figure 7-1-3-1-1•

During the sleeping period, heart rates recorded on passes

over tracking stations were generally as lov as 50 and

averaged between 55 and 60 beats per minute. However, when



p a g e 7 - 9

the pilot awoke and announced anything which was recorded on

the onboard recorder his heart rate immediately rose to about

80 which is the same value as during his working periods

earlier in the flight. After about 23:32:00 and for the

remainder of the normal orbital flight, the astronaut's

mean heart rate rose to a value of about 100 beats perminute. His first indication of a spacecraft system

malfunction occurred at about 28:34:00 when he noticed the

0.05g relay light had come on. Heart rate at this time rose

sharply to l48 beats per minute and then rapidly declined

to the low of 60 beats per minute and stabilized at a rate of

around 100 beats per minute. After a preliminary analysis of

the nature of the malfunction indicated by this 0.05g light,

the pilot's heart rate varied, with a peak of 142 beats per

minute while he was engaged in checking his ASCS system at

approximately 30:08:00. Again, the heart rate declined

rapidly to its resting level of approximately 100 beats per

minute.

At about 32:41:00, the pilot was advised to take 5 ing of

dextro amphetamine orally which he did very shortly after

receiving the advice. A gradual rise in the heart rate can

be seen, beginning at 33 hours elapsed time, with rather

marked swings in rate between levels as high as l4o beats

per minute and lows of about 80 beats per minute throughout

the remainder of the flight. The last significant inflight

change in heart rate occurred at retrofire when the heart rate

rose to l66 beats per minute for no longer than 20 seconds.

The heart rates during reentry varied between 120 and l4o beatsper minute until drogue parachute deployment when it spiked

to 184 beats per minute. It then gradually declined to

164 beats per minute when bioplug disconnect was accomplished

subsequent to main parachute deployment.

The changes in heart rate throughout this flight seem to

fall readily into two categories. The gradual increases in

rate with correspondingly gradual returns to normal resting

heart rate are seen in response to physical exertion. The

peak heart rate which was noted corresponded to levels which

would be expected following an equivalent amount of exertion

under Ig. A sharper rise of heart rate to high levels inexcess of l4o beats per minute is seen as a startle response

when the astronaut is evidently emotionally alerted to a

highly significant change in his environmental situation.

From an electrophysiology standpoint, the EGG was well

within normal physiologic limits during the major portion of

this flight. The A-V nodal beats noted during the prelaunch

period were nearly non-existent during the entire

C O M r i D D H T I A L



Page 7 - 1 0 ^ C O N F I D E N T I A L

3W hours of flight monitoring. A careful review of the

entire flight revealed that data from both leads of EGG

showed periodic changes in the character of the P wave and the

P-R interval which are consistent with a wandering pacemaker.

There were frequent prolonged sinus pauses during the flight

which generally are associated with deep inspiration by'the

pilot, and in every instance a sinus beat, rather than a

ventricular escape, followed the pause. One period when

this rule did not hold was during the sleeping time as the

astronaut was passing over the RKV tracking ship. At

17:10:00 and 18:45:00, the medical monitor reported a nodal

rhythm which was verified during the postflight examination

of the records. Figure 7-1-2.2-4 illustrates this variation.

Late in the flight, the sternal EGG lead became rather noisy

with a marked fluctuation of the baseline. This fluctuation

appeared at times to be synchronous with respiration and at

other times to bear little or no relationship to respiratory

movements. At this period in the flight, sinus arrhythmia

was somewhat more pronounced than it had been early in the

flight. A recurrent finding on the record consisted of a

simultaneous disruption of the sternal EGG recording with a

sharp negative impulse on the relatively insensitive respiratory

channel and a sinus pause showing on the side-to-side EGG

lead. It is tentatively believed that this characteristic

pattern resulted from either a habitual deep sighing breath

taken by the pilot or perhaps a repeated stretching motion

made in an attempt to relieve his cramped position.

Blood pressures did not vary remarkably during the flight

from preflight values, as shown in table 7.1.2.2-4.

Postflight analysis of the radiation film badges worn by

the astronaut revealed a total dose of from 15 to 20 millirads,

which is well below the maximum allowable dosage.

7.1.3.2 Clinical observations: The information contained in this

section was derived from reviewing the pilot's comments on

voice track of the onboard tape, conversations immediately

after the flight, and from his answers to a comprehensive

list of questions during the medical debriefing on the day

following recovery.

The noise and increased "g" forces associated with launch

were not uncomfortable and caused no problem. Specifically,

the pilot stated he felt very little vibration and had no

blurring of vision. He attributes this absence of blurring

to a slightly thicker' rubber pad which had been added to the

couch beneath his helmet. The cessation of powered flight

and the onset of weightlessness did not cause vertigo, tumbling



C O N F I D E N T I A L P a g e 7 - n

sensations, or other unusual symptoms. The pilot rapidly-

adjusted to zero g and described it as "pleasant, extremely

relaxing, and a real floating sensation." His observations

indicate that vision was entirely normal throughout the

flight.

He noticed a phenomenon of orientation similar to that

reported by the MA-7 pilot. He stated that after SECO and

during the first 20 minutes or so of gravity-free flight,

he felt the equipment kit located near his right arm was

rotated 90°• This phenomenon did not extend to the instrument

panel, window, instrument panel compartment or any other

structure within the spacecraft, and it was not troublesome.

The pilot's body was his reference for orientation to the

spacecraft and this relationship was never in question.

The astronaut stated that he did not feel particularly

hungry for the majority of the time during the flight and ate

primarily because it had been scheduled. However, later in

the flight he did feel hungry on one occasion and after

eating felt better. Because of problems with the food

containers and water nozzle during flight, he was unable

reconstitute properly the freeze-dehydrated food and could

only eat one-third of a package of beef pot roast. Therefore,

he subsisted on bite-sized cubed food and bite-sized peanut

butter "sandwiches." He avoided the bite-sized beef sandwiches,

since they had crumbled in their package. His caloric intake

during the flight was only 696 calories of the 2,369 calories

available to him at launch. He rapidly tired of the cubed

"snack-type" foods and this contributed to his low caloric

intake. Typical samples of the food types which were carried

aboard for the MA-9 flight are shown in figure 7-1-3.2-1.

The astronaut's water intake was also limited. When the

condensate transfer system would no longer permit fluid storage

in the main condensate bag (3.86 Ib) during the flight, he

was forced to put condensate water into one of the drinking-

water tanks before he had consumed all of its contents. Normal

operational procedures required the exclusion of condensate

water as a drinking-water source. He began drinking small

amounts from his survival kit water supply, as planned, but

wished to conserve this as much as possible. He was notreally thirsty until during the last orbital pass, but he

was so busy at this point he did not take time to drink.

Because condensate water was placed into the drinking-water

tank, in which an unknown amount of drinking water remained,

it is impossible to make a valid statement as to his water

intake during flight, but he did consume more than 1,500 cc.

The drinking water rapidly became unpalatable because of its

equilibration with 95° cabin temperature and an alteration of

U p



7 - 1 2

its taste by the plastic "bags in which it was carried. The

astronaut stated that he probably would have drunk more liquid,

including the dehydrated juices, if he had not experienced

difficulty with the mouthpiece of the drinking water tank.

At one point during the flight the pilot felt a vague gastricawareness or queasiness, but this feeling rapidly cleared when

he ate a little food and drank some water. At no time did

he have any nausea, vomiting, or other gastrointestinal problems.

He urinated without difficulty several times during flight

and stated that bladder sensations were normal. The urine

collection and transfer system worked well, and separate urine

samples were obtained at four different times during the flight.

It did require considerable time and effort to transfer the

urine to the storage bags, although not as much as was

encountered with the condensate system.

The astronaut had a very good sleep the night prior to launch

and was as rested as possible. He found, even early in the

flight, that when he had no tasks to perform and the space-

craft was oriented such that the earth was not in view from

the window, he easily dozed off for brief naps. There

were times when he awoke without realizing he had fallen

asleep. This dozing did not occur during times when there

were tasks to perform or items to see through the window.

During the period designated for sleep, he slept only in a

series of naps lasting no more than 1 hour each. His total

sleep time was about ^-5- hours. He awoke from these 30- to

60-minute naps feeling alert and rested, but 30 to ^5 minutes

later he would again doze off. He stated that if there had been

another person along to monitor the spacecraft, particularly

the ECS functions, he could have slept for much longer periods,

but still "no more than k to 6 hours in a day." Table 7-1-3-2-1

lists the estimated inflight sleep periods.

He had a brief period of confusion the first time or two

that he awoke, not realizing exactly where he was. However,

it took him only a very few seconds to become completely

awake and oriented. He reported that this brief period of

confusion did not occur later in the flight. The pilot

stated that he slept "perhaps a little more soundly" than

on earth and that he did dream, but he did not remember the

contents of the dreams. He did not notice significant changes

in dreaming from ground experience.

M. M.J\Li





p a g e 7 - i4 -CONFIDENTIAL

astronaut actuated the emergency oxygen flow rate for 30 seconds

which should have reduced the PCOp reading. It did not seem

to change the pilot's onboard reading noticeably, although

telemetry signals indicated a slight drop. At this time the

pilot closed his faceplate and felt that his respirations

were deeper and more rapid. This change in respiration could

not be confirmed by postflight examination of respiration

and heart .rate recordings. Although he felt more comfortable

with the faceplate open,, he kept it closed during the final

orbital pass and the reentry as planned. The PCOp gage

indicated about 5 mm Hg at reentry. This concentration is

not enough to cause symptoms of hypercapnia on the ground, and

there was no apparent interference with the pilot's normal

functions.

7-1.4 Postflight observations.-

7.1. .1 Recovery history: The spacecraft landed in the water about

4.4 miles from the recovery ship, the U.S.S. Kearsarge,

and was placed on deck approximately 40 minutes later. In

order to gain medical data as early as possible, the NASA

Flight Surgeon aboard the recovery ship was equipped with an

8-foot extension cord for the biomedical cable. Immediately

after the hatch was opened, this cord was attached to the

astronaut's biosensor plug and blood pressure fitting and

connected to the spacecraft onboard recorder in order to

record blood pressures and EGG before, during, and after

egress. This system was extremely effective in deriving

post-egress data.

The astronaut was then taken to the ship's sick bay where a

comprehensive medical examination and preliminary debriefing

were performed. The remainder of the debriefing was conducted

by the NASA Flight Surgeon in the admiral's in-port cabin.

The astronaut spent 48 hours onboard the ship. Details of

his activities during this 48-hour period are shown in

table 7.1.4.1-1.

7-1.4.2 Physical examinations: The postflight examination began

prior to egress from the spacecraft. Approximately 40 minutes

after landing, two measurements of the astronaut's blood

pressure were recorded while he was still lying in the space-

craft on the deck of the recovery ship. He was then able

to egress from the spacecraft without assistance and stand

erect on the deck vhile his blood pressure was again recorded

on the onboard tape. Later examination of this 3 minute

record shows that, while still in the spacecraft, the blood





p a g e 7 1 6 C O N F I D E N T I A L

it separated prior to this time although it appears probable

that it was loose, making partial contact and held by the

plastic insulation sleeve until the suit was removed. There

were some evidences of pressure on the skin at all lateral sensor

locations, but no signs of irritation by sensors, paste,

or tape. All sensors were securely in place and theelectrode paste seemed to have maintained its normal consistency.

At the sensor locations on the left lateral chest, there were

narrow semicircular marks that looked like a very shallow

cut with a sharp blade. These cuts may have been caused

by the thin edge of the tape where the rubber sensor disc

slightly overlapped it.

There were painful and slightly swollen red areas over each

patella caused by the pressure suit having been pulled

tightly across the anterior knee when the knee was flexed.

Other reddened areas were found over each posterior inferior

iliac spine and the posterior spinous process of the fifthlumbar vertebra. There was a diffuse redness over the

right lateral iliac area, but none on the left. No explanation

can be offered for this condition at the present time.

Additional findings of note were a bilateral conjunctivitis,

which probably resulted from drying of the eyes by the

constant oxygen flow and a slight reddening around the left

tympanic membrane. He complained that he had a little trouble

clearing his left ear during descent. Both ears "crackled"

for 6 to 8 hours after recovery as the oxygen in the middle

ear was gradually absorbed and replaced with air. This

condition is commonly seen in aviators when they have beenbreathing 100-percent oxygen.

Tilt table studies were performed at 1, 3, %-, and 19 hours

after landing. At no time did the astronaut have any

subjective complaints, nor were objective changes noted

except in pulse and blood pressure. Specifically, there were

no unusual color changes in the feet, as had been noted

following the MA-8 flight. The results of the tilt table

studies are tabulated and discussed under Special Studies.

The medical findings during the initial recovery examination

are shown in table 7.1.2.1-3,and included a blood pressure

of 90/80 mm Hg while supine, a heart rate of 86 beats per

minute, a respiration rate of l6 breaths per minute, a

body weight of 139r pounds, and a body temperature of

99.k° F taken orally. Three hours after landing, his urine

showed a specific gravity of 1.031,an(i the hematocrit

was ^9. These findings, combined with the clinical evaluation,

G O N r i D D F I T f



Page 7-17

indicate a moderate dehydration. As has been indicated

elsewhere, this dehydration resulted from a reduced intake

of food and water during the flight. Detailed results of

the blood and urine analyses currently available are contained

in tables 7.1.2.1-3 to 7.1.2.1-5. The reversal of the ratio

of lymphocytes to polymorphonuclear leukocytes during theweek following the flight, without a significant change in

the total count, is presently unexplained. This ratio has

since returned approximately to unity, but the study of this

phenomenon is continuing. A clinical electrocardiogram and

a chest X-ray completed the initial postflight examination.

The chest X-ray showed no changes when compared with that taken

before the flight on May 12, 1963. The EGG showed a moderate

rightward shift in the QRS and T axes when compared to that

of May 12, 1963. This shift is probably the result of a normal

cardiac position change.

The astronaut slept very soundly for 9 hours and awoke

cheerful and eager to complete the debriefing activities.

A brief examination the following day showed that the

conjunctival irritation, the hoarseness of his voice, most of

the skin pressure marks, and most of the evidence of dehydration

had disappeared. The areas of pressure over the knees were

still painful and somewhat more swollen than on the previous

day. The sharp semicircular marks were still much in evidence

and remained visible for several days.

Table 7.1.4.2-1 states the pilot's weight loss during several

preflight activities and the inflight experience. Intake and

output records for the first 2.khours after recovery indicate

a fluid intake of 3 900 cc and a urine output of 5^5 cc.

The pilot returned to the launch site on the fourth day following

launch and was examined the following morning. The same

medical specialists found him to be in excellent health.

The only changes noted were the persistent slight erythema

and tenderness of both patellae, resulting from the pressure

areas in the suit, a continued, rightward shift in the QRS

and T axes of the EGG, and a persistence of the previously

noted alteration in blood count. The EGG- shift had become

less apparent, however. The laboratory studies of blood and

urine are contained in tables 7-1-2.1-3 to 2.1.2.1-5.

The pilot remained in good health and maintained his high morale

following this examination. He participated in debriefing

sessions and other postflight activities without further medical

change.

lOOMilDBMTAL



Page 7-18

7.1.5 Special studies.-

7-1-5-1 Tilt test evaluation: The medical examination performed

immediately after the MA-8 recovery suggested an alteration

in the pilot's cardiovascular responses to position changes.

In order to obtain more quantitative measurements of theseresponses, an operational tilt procedure was developed for

shipboard use. This procedure utilized a Stokes' Litter with

cross-bars added for lifting and stabilization. These modi-

fications permitted a tilt of 70° from the horizontal in

3 to 4 seconds. The individual being tested was comfortably

secured in the litter, without circulatory interference, by

straps across the knees and the upper chest.

Heart rate and blood pressure measurements were taken at

least every minute on all tests and were chosen as the primary

indicators of altered function, in conjunction with observation

of visible reactions and subjective comments. Operational usecalled for minute heart rates calculated from 15-second counts

of the right radial pulse with clinical blood pressures taken

from the left arm. Greater capability in the Space Medicine

Laboratory in Hangar S permitted simultaneous determination of

both clinical and BFMS blood pressures and continuous recording

of respiration rate and EGG from the biosensor system. Minute

heart rates were determined from the directly recorded biosensor

data by using 12-second counts made every 30 seconds.

Minute respiration rates were determined from 30-second counts

made each minute. There were no apparent differences between

the clinical and biosensor values, but continuous EGG readingsproduced interesting additional information.

The procedure was carried out in the following manner. After

k sets of similar control values, the individual was tilted for

5 minutes and values were sampled at least every minute, then

returned to the horizontal position for a recording of at least

4 more sets of similar values. Thus, the minimum time for the

complete test was 13 minutes. In order to superimpose a further

cardiovascular stress, Flack Tests were used in some of the tilts.

This test utilizes a tube with an orifice through which the indi-

vidual exhales after a maximum inspiration, producing a constant

pulmonary overpressure of ^0 mm Hg. The Flack Test lasted

15 seconds and was conducted from 3;j to minutes after the

individual was tilted to the 70° position.

Preflight results were obtained from 11 tilt tests on the flight

astronaut from January 5 to May 10, 1963. Flack Tests were per-

formed with k of the tilts. All of these tilts were performed

in conjunction with a spacecraft checkout procedure which required



Page 7-19

at least 2 hours in the spacecraft couch In the semisupine

position. The time between the prerun tilts and the procedure

varied from 1 to 5 hours because of uncontrollable operational

factors. In each case, the postrun tilts were conducted from

5 to 15 minutes after the procedure, and on January 5, 1963,

a second postrun tilt was performed 1 hour after the first.

The heart rate and blood pressure values are summarized in

table 7.1.5-1-1 and illustrated in figures 7.1.5.1-1 and

7.1.5.1-2. The preflight results fall within the ranges

reported in the literature. In the prerun period, most heart

rates were between 55 and 80 beats per minute. The tilt pro-

duced a rise in heart rate varying from 5 to about 20 beats

per minute within 30 seconds. This reading gradually increased

during the first 2 minutes to rates of 80 and 9° beats per

minute, at which point it stabilized. Post-tilt values between

100 and 110 beats per minute occurred after a 6 hour run, which

was more than twice as long as any of the other runs.

At the beginning of the Flack Test, a bradycardia for 3 or 4

beats usually occurred, followed by an increase in rate to 80

to 90 beats per minute. On several occasions, the maximum

observed rates of 110 beats per minute followed a Flack Test.

The sudden release of the increased intrathoracic pressure

again produced a transient bradycardia followed by an "over-

shoot" of 10 to 15 beats per minute. Conclusion of the tilt

period consistently produced an immediate drop in rate to the

pretilt range. Respiration rates were without significant

change and are not reported. The increases in diastolic blood

pressure were the most' remarkable produced by the tilt. The

mean increase was 15 mm Hg, but many of the diastolic pressures

rose 20 to 30 mm Hg. An initial systolic drop was followed by

a compensatory rise. Post-run tilts produced somewhat more

striking blood pressure changes, with narrowing of some pulse

pressures to as little as 6 mm Hg. The maximum systolic levels

followed Flack Tests, without an assosicated diastolic change

of significance.

The EGG demonstrated expected alteration of the QRS axis

secondary to position change. Decrease in size of the QRS

was especially prominent in the chest lead as a consequence

of E wave depression. There were sinus pauses with an occa-

sional aberrant complex of ventricular origin. The usual pre-

tilt sinus arrhythmia disappeared with the rate increases. The

Flack Test produced dropped beats and occasional premature ven-

tricular contractions during the period after sudden release.

C O M r i D E M T I A L





Page 7-21

Tilt test

no.

PretiltMean I Minimum I Maximum

Tilt

Mean I Minimum

Preflight pulse

29 17 4o _ 21

Postflight pulse

1 to 3

4

25

37

10

36

38

38

17

26

Maximum

Posttilt

Meanl Minimuml Maximum

pressure, mm Hg

6 58 31

pressure, mm Hg

6

14

42

46

29

46

10

20

40

46

36

50

7.1.5.2

The blood pressure responses to the final tilt were nearly

normal, but still showed a delayed compensation for the systolic

drop. Wo visible objective changes occurred and there were no

subjective symptoms.

In summary, the preflight tilt tests produced expected

cardiovascular compensatory reactions insofar as they could

be demonstrated by heart rate, blood pressure, and EGG data,

and all of these tests were well tolerated. The postflight

tilt tests demonstrated the presence of moderate orthostic

hypotension, with far greater heart rates required to maintain

effective cardiovascular function. Compensation was achieved,

however, and the pilot did not even develop near-syncope.

Tilt studies of responses after stresses similar to those

experienced during flight are not available, and inadequate

water intake and weight loss create difficulty in interpretingthese results. It is well known that dehydration and decreased

blood volume are some of the factors which will decrease

cardiovascular stress tolerance.

Calibrated work: A device for calibrated work consisting of

a short plastic handle and expandable bungee cords(see

fig. 7-1-5-2-1) was fixed within the spacecraft near the

astronaut's feet. A limiting cable ensured repeatability

of handle travel, requiring 65 pounds of force for each full

extension. At 2:25:00 and again at 7:4l:00, the astronaut

recorded his blood pressure, pulled the device 30 times in as

near 30 seconds as possible, and again recorded his blood

pressure. The results of these two work periods were compared

with 5 such periods performed at normal gravity in the space-

craft and in the procedures trainer.

Subjectively the astronaut could tell little difference

between the work performed under normal gravity and under

zero gravity, the effort under zero gravity being, if anything,




slightly easier. During flight he felt his post-workbreathing was not as labored and he thought his heart rate

returned to pre-work values more rapidly. The data aresomewhat, affected by the difficulty he had in freeing thework handle from its restraining clip during flight.

Analysis of the data does not show any striking differencesbetween the one gravity and zero gravity work periods.Inflight mean heart rates during the calibrated work periodare 2.6 beats per minute higher than preflight, but hisinflight mean heart rate before work is 15 beats per minute-

higher; therefore this difference is not though to ~vbe significant,

The results are given in table 7-1-3-1-1 and presentedgraphically in figure 7-1-5.2-2. One preflight heart rateduring work was l6o beats per minute. This value occurredat the only time in one of the seven periods that he workedover 0.7 minute and probably reflects the prolongation of

the work period rather than indicating a higher work load.During the 18-second recovery period after the test, thepreflight mean heart rate is seen to drop to 11 beats perminute over the preflight value, while during the flight itfalls to 17 beats per minute over the prework mean. Thisdifference is also not thought to be significant. Bloodpressure readings taken before and after work show nosignificant changes from baseline values.

In summary, based on the inadequate evidence obtained duringthe MA-9mission, it appears that for brief periods of workfollowing relatively short periods of zero gravity, the

cardiovascular system shows no significant changes in heartrate or blood pressure. Subjectively,.calibrated workunder zero gravity seems easier, as do most tasks requiringphysical effort.

7.1.5.3 Special clinical studies: Retinal photography, urine andplasma electrolyte determinations, and plasma enzyme studiescomprise the special clinical studies. The retinal photographs,taken after the flight for comparison with preflight pictures,are not ready at the present time. The preliminary resultsof the plasma electrolyte determinations are available

and appear in table 7-1'5-3-1- If the increase in serum

calcium is valid, it is a small change. The results of theurine-electrolyte determination are presented in table 7.1.2.1-5;

however, the plasma-enzyme determinations are not yet available.

7-1.6 Conclusions.-1. There was no evidence of significant degradation of

pilot function directly attributable to the space flight.Thirty-four hours of zero gravity were well tolerated andall body functions appeared unaffected during flight.





Page 7-24

TABLE 7.1.2.1-1.- PILOT PREFLIGHT ACTIVITIES

[Selected activities for which medical study or support was performed!

Date Activity Medical study or support

5

22-23

23

7

8

10

11

12

15

Altitude-chamber

spacecraft checkout

Hangar flight

simulation

Flight simulation no. 1

T-10 day physical

examination

Mission simulation

(procedures trainer)

Launch simulation

Flight simulation no. 3

T-2 day physical

examination

Countdown(flight canceled)

Flight countdown

Physical examination before and after

Background data (biosensors)



Low residue diet (3 days) and flight food (2 days)

Physical examination

Background data (biosensors)Timed urine collection

Physical examination, ^5 minutes



Timed urine collection




Begin controlled dietBlood specimen, 50 cc




Begin low residue diet

Comprehensive medical examination, 2 hours

Blood (30 cc) and urine specimen

Physical examination before and afterTimed urine collection

Blood specimen, 30 cc

Physical examination

Aeromedical countdown

Awaken 2:51 a.m. e.s.t.

Launch Q:0k a.m. e.s.t.

CONriDENTlAr



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TABLE 7.1.3-2-1.- INFLIGHT SLEEP PERIODS

[other unrecorded naps occurred!

Time, c.e.t.

02:10:15 to 02:11*: 00

05:1*0: 00 to 05: 5:00

13:50:00 to Ik:k6:00

Ik:20: 00 to lk:kj:00

15:11:00

15:20:00 to 16:05:00

16:28:11

16:50:00 to 17:50:00

18:20:00 to 18:25:00

18:1*0:00 to 19:27:00

19=38:39

21:22:kk

27:26:08

Estimated duration,

min

k

5

56

27

(a)

^5

(a)

60

5

7

(a)

(a)

(a)

Source

Onboard tape

Astronaut reco

Onboard tape

Astronaut reco

Onboard tape

Astronaut reco

Onboard tape

Astronaut reco

Astronaut reco

Astronaut reco

Onboard tape

Onboard tape

Onboard tape

Total sleep recorded: k hours and 9 minutes

Short naps, duration not determined-



Page 7- 3 8 C O N F I D E N T I A L

TABLE 7.I.IK 1-1.- PILOT POSTFLIGHT ACTIVITIES

Date,

1963

Time, local Midway

(a)

Activity

May l6

May 17

May 18

May 20

12:25 P-m.

12:55 p.m.

1:09 p.m.

1:12 p.m.

1:15 p.m.

l:U5 p.m.

J:00 p.m.

3:30 p.m.

3:U2 p.m.

U:10 p.m.

5:^5 p.m.

7:11 p.m.

9:30 p.m.

7:00 a.m.

7:1+0 a.m.

8:00 a.m.

9:00 to 11:00 a.m.

2:00 to 5:00 p.m.

7:00 to 9:00 p.m.

1:00 p.m.

9:00 a.m. e. s. t.

Landing

Spacecraft on deck

Blood pressure, recumbent in spacecraft

Egress and blood pressure standing

Physical examination begun in recovery ship sick bay

First tilt table procedure

Examination completed

First postflight urination

Second tilt table procedure

First postflight meal

First postflight bowel movement

Third tilt table procedure

To bed

Awakened

Fourth tilt table procedure and brief medical

examination

Breakfast

Self-debriefing

Technical debriefing

Medical debriefing

Left recovery ship

Comprehensive postflight medical examination at

Patrick Air Force Base, Florida

To convert times to e.s.t., add 6 hours



Page 7-39

TABLE 7.1. .2-1.- RECORD OF PILOT'S WEIGHT CHANGES

During the J-êek period prior to flight, the pilot's maximum weight was

1 9r lb and his minimum weight was 1^6 l"b. His weight on launch morning

!

was 1 7 lb and his weight on the recovery ship was 139r It.

Date Activity Duration, hr Weight loss, lb

Pre flight

January 5, 19 3

April 23, 1963

May 8, 1963

May 10, 1963

May 14, 1963

Altitude -chamber spacecraft

checkout procedure

Flight simulation

Launch simulation

Flight simulation

Canceled launch

9

7

8|

6

8

3-5

2.0

3.0

2.0

1-3

Flight

May 15/16, 1963 Orbital flight ^ 7-75

C O M T D D F r i T A L



Page 7 - 1+0

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Page 7 - 4l

TABLE 7.1.5.3-1.- BLOOD CHEMISTRIES

JA11 "values must be reverified

Determi-

nation

ml

hosphorous,

ml

March 12,

1963

k.YI

105

6.0

4.2

153

k.6

May 8,

1963

4.28

106

6 - 3

3 - 5

151

4.6

May 12,

1963

al|.60

100

6.0

u . u

161

5.1*

May 1 ,

1963

U.22

104

6 .6

k . h

14U

5 - 2

May 16, 1963

landing + 2

hours

4.67

104

6 - 3

4.5

153

5 - 2

May 17, 1963

Landing + 24

hours

4.56

102

6.2

4.0

147

5-0

May 20,

1963

4.22

104

6.2

3-4

146

4.9

This value is particularly in question.



p a g e T - ^2 tJONFIDENTIAL

7.2 Pilot Performance

7.2.1 Preflight training.-

7.2.1.1 Spacecraft checkout activities: Intensive participation in

the preflight checkout activities enabled the pilot to be-come familiar with the MA-9 spacecraft and launch-vehicle

systems. Participation in these activities was particularly

important because the pilot had the opportunity to operate

and to become familiar with the characteristics of the modi-

fied systems and switching procedures peculiar to his own

spacecraft. Table 7*2.1.1-1 summarizes the checkout activi-

ties in which he spent 73 hours and 50 minutes in the space-

craft. This table does not include the many hours spent

by the pilot in preparation, troubleshooting, monitoring

tests as an outside observer, and posttest analysis. In

addition, as backup to Astronaut Schirra, the pilot spent

approximately ^0 hours in the MA-8 spacecraft, which addedto his knowledge of the Mercury spacecraft in general.

7.2.1.2 Flight simulation training activities: Table 7.2.1.2-1

summarizes the training activities on the Cape Canaveral

procedures trainer from March 15 to May 13, 1963* The pilot

spent 33 hours and 30 minutes on the trainer accomplishing

^3 turnaround maneuvers, 62 simulated manual retrofires,

8l simulated systems failures, and 28 manual reentries. In

addition, the pilot spent approximately 20 hours on thepro-

cedures trainer during the MA-8 preflight period.

The pilot spent the majority of his time during these sessions

on the detection and circumvention of simulated systems failures

and various mission anomalies which would require an abort

during the launch phase of the flight. The pilot concentrated

on these areas because of their critical importance and because

the procedures trainer is best equipped to simulate these

phases of the mission. He did,however, spend approximately

one-third of his time during these training sessions on the

normal flight activities specified in the flight plan and

practicing of retrofire and reentry attitude and rate control

maneuvers. The pilot also participated in several launch

abort and network simulations during which mission rules and

operational procedures were extensively used and reviewed.

7.2.1.3 Flying proficiency: Throughout the preflight training period,

the pilot continued to maintain flying proficiency in high-

performance aircraft. He logged 6k hours and 35 minutes in

flying time from January 1 to May 10, 1963 (see table 7.2.1.3-1),

Flying high-performance aircraft is an important complement to

the more direct space-flight preparation activities on the

static trainers, because it requires the pilot to maintain the

COMTDONTIAlr



Page

ability to make quick and accurate analyses and decisions under

actual operational conditions.

7.2.1.U Systems and operational briefings: The pilot received two

series of formal systems briefings which were oriented as

much as possible toward the operational requirements of themission. The first series of briefings required 3 days time

during which each spacecraft subsystem was extensively re-

viewed, while the second series required only 1 day since

its purpose was to review recent spacecraft systems

modifications.

The pilot received two series of briefings on inflight

experiments, each series requiring 1 day to complete. During

these presentations'1the experiment, its purpose, its associated

equipment, and the operational procedures were discussed, final-

ized, and integrated into the flight plan. In addition, the

pilot spent many more hours with various systems and operationspersonnel on an informal basis in order to establish optimal

operational procedures.

7.2.1.5 Preflight-operations schedule: Table 7-2.1.5-1 summarizes

the major activities in which the pilot was involved from

January 2, 1963, until launch. In addition to the large

amount of time spent in briefings, spacecraft checkout

activities, flying, and on the procedures trainers, the

pilot was also able to take part in several special training

activities. Among these activities were four sessions at the

Morehead Planetarium, during which he reviewed the celestial

sphere and participated in a simulation of the flashing light

experiment; acceleration refamiliarization in the Johnsville

centrifuge; and recovery and egress training and survival-

equipment exercises.

7-2.1.6 Training analysis: The pilot, like his predecessor, con-

centrated his major effort in learning the spacecraft' s

systems and their operational characteristics. The pilot

achieved a high level of skill on the procedures trainer

in coping with systems failures and operational anomalies.

Particular concentration was placed in developing proce-

dures for accomplishing inflight activities with a minimal

expenditure of onboard consumables. The pilot practiced

these activities on the trainers only after he knew each

system or operation extremely well and all spacecraft sys-

tem modifications had been completed.

He also obtained considerable practice in performing man-

ually such tasks as the turnaround, retrofire, and reentry

maneuvers by using the attitude and rate indicators. A good

understanding of the spacecraft's control systems was provided




p a g e 7 - * C O N F I D E N T I A L

through systems briefing and the use of the transparent

gyro simulatior; and the yaw-recognition trainer helped

him to prepare for those inflight activities, such as

gyro realinement and- flight maneuvering with the use of

external references, that are not properly simulated on

the Cape Canaveral procedures trainer. The pilot reportedthat the'one inflight experience for which he was not well

prepared was the out-the-window view. Since this view can-

not be easily simulated, it was a novel experience to the

pilot and did cause him some distraction, but i t was not

a major problem.

Because of experience gained from past Mercury flights, a

fairly long preflight period, and a diligent and intensive

training program, the pilot reported, and the results verify,

that>he was very well prepared for this flight. He reported

that, in general, there was a proper proportion of training

effort placed upon the different aspects of his mission. The

only significant areas of pilot preflight preparation that the

pilot believed could be improved or supplemented were:

1. An earlier finalization of systems and attendant

operational procedures

2. Additional improved external-reference simulations

3- Additional training on the recognition and observa-

tion of celestial bodies

k. Less egress training through the top hatch

5. Somewhat more concentration on normal as opposed

to emergency procedures during the simulation of the launch

phase of the mission on the procedures trainer

7.2.2 Flight-plan activities.-

7-2.2.1 Flight-plan description and results: The flight plan was

designed so that the operational test requirements and

approved experiments could be accomplished within a 22-pass

mission. The duration of the mission required the scheduling

of an 8-hour sleep-rest period.

The mission was separated into four phases with go-no-go

decisions at the end of the 1st, 7"th,and l6th orbital passes.

In order to conserve consumables and still accomplish all of

the operational and experimental objectives, the pilot was

to refrain from using the automatic stabilization and control

system (ASCS) bus and allow the spacecraft to drift for long

periods of time. The turnaround maneuver was scheduled to be



Page 7 -

accomplished manually by the pilot; however, the retrofireand reentry events were to be controlled by the automaticcontrol system.

The pilot's adherence to the flight plan was excellent, andalmost all of the flight activities were completed as planned.

Table 7.2.2.1-1 presents a summary of the major events thatoccurred during the flight. Up to the time at which a fail-ure involving the automatic stabilization and control system0.05g relay circuit and the subsequent loss of all ASCSa-c power occurred, all of the major planned events wereaccomplished at approximately the programed times. Thereafter,the pilot concentrated primarily on checking out the status ofsystems and isolating the malfunctions that had occurred. How-ever, he still managed to complete most of the flight-planactivities that were scheduled after the malfunctions werenoted.

The first phase of the mission was accomplished as planned andall systems were thoroughly checked out with the exception ofthe fly-by-wire (PBW)high thrusters. A decision was made toproceed into the second phase of the mission at the end of the1st orbital pass.

The second phase was accomplished as planned. The flashinglight was deployed at 3:25:38 and shortly thereafter theASCS a-c bus was powered down. Although the pilot did notobserve the flashing beacon during the first night phaseafter deployment, he was able to make several observations

during the next two night phases. The telemetry mode ofthe TV transmitter was checked over Hawaii and Californiaduring the 3^1 orbital pass, and the pilot sent continuous-wave code signals to Bermuda during the ^th orbital pass.At 6:22:00, the pilot turned the cabin fan and cabin-coolantflow off, powered up the ASCS a-c bus, and returned the space-craft to ASCS control. At about 8:21:00, the pilot returnedto FBW-low and pitched down for observation of the groundlight. After successfully completing this experiment, thespacecraft was returned to ASCS control. At about 09:00:00

the pilot returned to FEW-low and at 9:01:00 and 9:07:00,

he attempted without success to deploy the tethered balloon.

He then returned the spacecraft to ASCS control and continuedto check out the spacecraft systems in preparation for thedecision to proceed with the third phase of the flight.

A go decision was made at 10:00:00, and the ASCS a-c bus

was powered down at 10:26:36. Shortly before the startof the rest period, the pilot took some general purposephotographs, had a meal, and checked the manualpro-portional and FEW-low control modes. His sleep-rest period



lasted from about 1J:50:00 to 21:23:00, at which time the

pilot reported to Muchea. At approximately 23:30:00, the

pilot powered up the ASCS a-c bus, checked manual pro-

portional and then alined the spacecraft by using FBW-low.

He then went to ASCS control for another checkout of the

systems and to prepare for the dim-light-phenomena photo-

graphs. This experiment, as well as the horizon-definition

quadrant photographs, was completed on schedule.

The systems were checked prior to the decision to proceed

into the fourth and final phase of the flight, and the ASCS

a-c bus was powered down at about 25:17:00- The pilot con-

tinued with the flight plan and completed moon-earth limb

and the infrared weather photography as scheduled.

At 28:3 :3 ; "the0.05g relay circuit latched in; however,

the pilot continued with the flight plan with the exception

of those activities requiring ASCS control. He performedthe first half of the HF antenna test on schedule. The ASCS

a-c bus was powered up twice after the 0.05g light illumin-

ated and the pilot made checks which verified that the attitude

gyros had been lost. However, these checks revealed that the

auxiliary damping mode was functioning properly and that ASCS

could be used during reentry. As a result of the loss of the

attitude gyros, the pilot did not take the horizon-definition

photographs.

The cabin fan and the cabin-coolant flow were turned back on

at 32:05:00, and the pilot stowed the onboard equipment shortly

thereafter. At about 33:07:00 planned use of the ASCS for re-entry was abandoned when it became evident that all ASCS a-c power

had been lost. The pilot did not check out the FBW-high

thrusters prior to retrofire as ASCS problems were receiving

his primary attention during this period. The retrofire,

retropackage jettison, and reentry events were all initiated

and controlled manually by the pilot. He momentarily checked

his FBW-high thrusters subsequent to retropackage jettison;

but he was unable to detect proper high-thruster action and there-

fore elected to control the spacecraft, during reentry by

using FEW, high and low, and manual proportional simultaneously.

The pilot was able to control the reentry manually by using this

"double-authority" control, and the remainder of the descentwas normal.

The pilot managed the operation of the telemetry system, the

C-band beacon, and the S-band beacon throughout the mission

according to the scheduled program. He completed all but

three of the short status and consumable reports which were



Page ? -

programed for approximately once per orbital pass except

during the rest period, and he completed all three of theplanned long status reports.

7.2.2.2 Operational equipment: The pilot was provided with food,drinking water, and equipment designed to obtain quantitative

data during experiments, operational information in casecommunications were lost, and information which might aidhim in completing normal and emergency operating procedures.These items were stored in three locations: a special equip-ment storage kit to the right of the pilot's shoulder, aspecially shaped container with a writing-desk lid that couldbe pushed out of the way or pulled up into the pilot's lap,and the instrument panel storage compartment.

The space in the equipment kit was devoted entirely to experi-mental equipment and food. (See fig. 7-2.2.2-1.) As inprevious manned orbital flights, the pilot reported extreme

difficulty in locating, acquiring, and stowing items in thiscontainer. Because of the limited space available, a moresuitable location has not yet been found.

Equipment stowed in the writing desk (see fig. 7-2.2.2-2)

was accessible, with the exception of the small standardlight source which could not be dislodged from its velcroattachment point between two of the desk's structural ribs.In addition, the desk's position when extended was too close

to the pilot's lap during weightless flight. The desk containedequipment for experiments and observations, food, and two itemsof considerable operational value to the pilot, the navi-

gation booklet and the star navigation charts. The naviga-tion booklet contained a map in two sections, each coveringthe orbital ground track of 11 passes with ground-elapsed-timemarks at 1-minute intervals. The map sections included informa-tion relating to primary- and contingency-recovery andtracking-station locations. The booklet also contained theworld-wide weather forecast map; the nominal and "red-line"curves for automatic and manual fuel usage, oxygen consumption,and recording-tape expenditure; a table of nominal retrosequencetimes; normal, emergency, and experiment checklists; and aspecial list of abbreviated continuous-wave Morse code signals.

The star navigation charts were so constructed that a slidemechanism could be positioned by timing marks to provide astar-field picture equivalent to the window view with thespacecraft at attitudes of 0° in roll and yaw and from 0° to

-3 -° in pitch.





CNTIAL Page 7 -

therefore, he chose to observe and photograph the launch-vehiclesustainer stage prior to assuming orbit attitude.

The pilot reported that the maneuver felt just as it had onthe procedures trainer and he used the rate and attitude indi-

cators in accomplishing it. The maneuver required 0.2 poundof control fuel. This value is approximately 5 percent ofthe control fuel typically required by the ASCS during an

automatic turnaround maneuver.

7.2.3.2 Gyro realinement maneuvers: The gyros were realined withthe spacecraft on three different occasions. Table 7-2.3.2-1gives the time required and the fuel used for each of thesemaneuvers. The first two maneuvers were performed by usingFBW-low; however, the third realinement was accomplished inpart with FBW-low and the remainder with manual proportional.In the two cases in which data are available, there were virtu-ally no errors between the scanner and gyro readouts in pitchand roll when the scanner first came on the line. The dataalso indicate that yaw was in good alinement at the completionof both maneuvers. During the alinement periods, the space-craft rates were held to l°/sec or less, and this reducedrate probably accounts in part for the small amount of auto-matic-system fuel which was used. All three realinementmaneuvers were accomplished on the day side of the orbit;however, the astronaut performed an equivalent maneuver onthe last night phase of the mission, apparently with goodresults. The results indicate that realinement of the gyros

to the true spacecraft attitude is easy on the day side. Re-alinement on the night side is not difficult but requires a

longer time period than on the day side. That the pilot wasable to accomplish these maneuvers so accurately and with solittle fuel usage can partially be attributed to the factthat the spacecraft gyros, for the first time, could be re-alined while at the -3^° position in pitch.

7.2.3.3 Drifting flight: The pilot spent approximately 4 3 percentof the orbital phase in attitude-free drifting flight withthe gyros caged and with no control-system operation. He also

spent another 39 percent of the flight with the gyros caged,

but he used either the manual proportional or the FBW-low con-trol mode at infrequent intervals. The pilot reported duringthe flight and in subsequent postflight debriefings thatdrifting flight presented no problems either with respectto his orientation or use of the spacecraft systems. He

reported that drifting flight was very pleasant, and therealization that he was conserving electrical power and controlfuel was comforting to him. He also commented that, during



p a g e 7 - 5 0

drifting flight, there were long periods with no observable

spacecraft rates on the indicators and also periods when rates

1°of -^/sec were present. The highest attitude rates during

drifting flight seemed to predominate in roll, but did alter-

nate between axes.

7.2.3.4 Yaw-attitude determination: Determination of yaw attitudes

and rates in daylight was reported to be quite easily accom-

plished, even when only a small portion of the earth was in

view through the window. The pilot felt confident that he

could accurately aline the spacecraft directly towards or

away from his direction of motion over the ground within 1°.

At the 90° position, that is, perpendicular to the direction

of travel, he believed that his accuracy might be degraded to

±10°. He therefore chose to use his attitude indicators when-

ever a precise 90° Yaw maneuver was required. The instrumenta-

tion system does not indicate yaw attitude directly; however,

analyses of horizon-scanner outputs after each uncaging maneuver

and return to ASCS orbit mode suggests that the pilot had deter-

mined the proper yaw attitude to within 5°. The pilot used

several visual cues to determine yaw attitudes and rates dur-

ing daylight, such as the "streaming" by of terrain features,

cloud patterns, or both, the convergence point of these flow

lines, and the tracking of terrestrial objects or cloud promi-

nences across the window. The view through the window kept

the pilot constantly aware of his rapid motion over the ground,

and he reported absolutely no difficulty in orienting quickly

to retroattitude.

The pilot reported that yaw-attitude determination at night

was not difficult, although it usually required a longer period

of time, particularly in the absence of good external visual

cues. The determination of yaw attitude at night was accom-

plished in two different ways. When the moon was illuminating

the earth and if the pilot was sufficiently dark-adapted, he

used the motion of terrestrial and cloud features to find the

points directly approaching or receding along the spacecraft's

ground track, just as in the daytime. In some cases lighted

cities or the glow of their illumination through thin cloud

decks provided a good reference for observing their direction

of relative motion, even without moonlight, for yaw attitude

determination at night.

The second method of determining yaw attitude at night was used

during periods when ground objects were not visible. In this

situation, the star field was the pilot's only source of in-

formation for finding yaw attitude," and more time was required



Page 7-51

to find identifiable stars and constellations. He found the

star charts to "be invaluable for identification and for re-

establishing his yaw attitude by using stars near the planeof the spacecraft's orbit. Star recognition was complicated

by the restrictive field of viev through the spacecraft window,but prominent groupings of stars were periodically available

during the early part of the night phase before moonrise.

The pilot's inflight reports and postflight comments of starsin view during the first night after deployment of the flashinglight indicated that he may have been some 30° to the right ofhis intended position while searching for the light. However,this was his first attempt to aline l80° in yaw at night andhad the added complication of the reversal of the normal

orbital attitude shown on the star charts. He accreditedhis prompt observation of the beacon just at sunset of thesubsequent night phase to the easier task of accuratelyalining the spacecraft during daylight towards the beacon's

expected position. It should be noted that the pilot expertlyperformed his most critical night yaw alinement, that for retro-fire, without attitude indications by using star and groundreferences only.

A convenient method of yaw determination was put forward

by the pilot after observing the relative motion of the so-

called fireflies, which he and pilots of all previous orbital

missions have reported as having seen. The luminous particles,

which appeared to emanate from the thrusters, were observed

to move outward from the spacecraft, then to recede back along

the spacecraft's trajectory in the manner of a contrail, re-maining visible for many seconds. The pilot suggested that

by positioning the spacecraft relative to the motion of these

particles, an accurate determination of the 0° pitch positions

might be possible.

Retrofire: It was originally planned to have the pilot use

ASCS control during the retrofire period, with the manual

proportional control system ready as a backup, if necessary.

As a result of the loss of ASCS power, the pilot was required

to initiate manually the retrofire event and to control the

spacecraft during the retrofire period by using the rate gyro

indicators and the view of the earth through the window as hisattitude references.

The pilot, realizing that he would be conducting the retro-

fire maneuver shortly after sunrise of the final daylight

phase, oriented and maintained the spacecraft near retro-

attitude throughout the last night period. He made very

small attitude corrections by using stars and clouds as





ICNTTCAL Page 7 - 53

7.2.3-6 Reentry: Because of the loss of the ASCS power, the pilot

was required to position the spacecraft manually to the proper

reentry attitude by using external visual reference to insert

manually the proper roll rate, and to damp manually the reentry

rate oscillations. The pilot checked his FEW control mode

shortly before the nominal time for the 0.05g event and was

not satisfied that the high thrusters were working properly.He, therefore, elected to control reentry rate oscillations

by using both the manual proportional and the FEW control

systems; thereby, U9 pounds of thrust capability was made

available to him about the pitch and yaw axes.

The pilot maintained the earth horizon in the window for

attitude reference until shortly before the nominal time for

0.05g. He then allowed the spacecraft to pitch up slowly to

reentry attitude and reduced his angular rates to near zero.

During the early portion of the reentry, he was easily able

to damp the small and rather slow oscillations by using FBW-low

thrusters and the manual proportional control mode. At approxi-

mately 1 minute and 30 seconds prior to peak reentry deceler-

ation, the pilot inadvertently actuated the IW-high yaw thruster

This actuation resulted in almost ^9 pounds of thrusting and

added to the amplitude of the oscillations. He brought them

back under good control within a 1-minute period and main-

tained positive control of the oscillations through drogue

parachute deployment. The pilot had no further difficulties

in controlling the reentry oscillations except for a brief

period during maximum deceleration. He was unable to manipu-

late the control handle at this time because the g-forces

pulled his arm away from the control handle and into a trough

on the arm rest.

The maximum frequency of oscillation occurred at peak deceler-

ation and lasted on the order of 0.9 second. Maximum rates

were approximately ±159/sec or 20°/sec, with a maximum ampli-

tude of approximately ±10° in both pitch and yaw which occurred

just after the peak deceleration period. The pilot reported

that he believed he needed dual authority control to be effective

during this reentry period.

7.2.k Systems management and operational procedures.-

7-2. .1 Control system utilization and switching: Table 7.2.U.1-1

and table 7-2.2.2-1 show control-mode usage and switching

during the flight. In general, control system usage was al-

most identical to the planned rate until the 0.05g relay pre-

maturely latched, with a subsequent loss of ASCS power. The

pilot was able to perform even more maneuvering during the



Page

period than scheduled "because of this effective and frugal

use of control fuel. Following the failure of the amplifier-

calibrator, the pilot was required to deviate somewhat from

the planned control-mode schedule, and for the remainder of

the mission manual proportional or FEW was used for space-

craft control.

The pilot was highly successful in switching from the manual

proportional and FEW to ASCS control. The orientation high-

thruster mode was never inadvertently actuated. The pilot

alined the spacecraft manually to the proper orbital attitude

on eight occasions during the flight. The spacecraft always

passed through orientation low before dropping into orbit

mode. The -maximum excursion that was evident during periods

of switching from manual to automatic control was 5° in atti-

1°tude and ~z /sec in rate.

The pilot's utilization of the manual proportional control

mode was much better than what was expected, based on the

previous flight results. By making very rapid hand-controller

motions in this mode, he was able to produce a thrust level

which was much less than the expected level of approximately

^ pounds. The performance of the manual proportional control

system, by using this technique, compared very favorably with

the performance of the FB¥-low system both with respect to

fuel usage and fine attitude control. The manual proportional

control mode was utilized during the last orbital pass for

alinement and attitude hold prior to retrofire, during which

slightly more than 2 pounds of manual fuel was expended.

The pilot did not at any time inadvertently use double author-

ity during the mission. The only time double authority was

used was for damping the reentry oscillations. This choice

was made because the pilot was not satisfied that the FBW-high

thrusters were working properly, and he therefore elected to

use FB¥ as a backup to the manual proportional control system.

7-2.h-,2 Conservation on consumables: The consumables of which the

pilot could directly affect the usage were control system fuel,

electrical power, and onboard tape-recording capability. Since

the use of these consumables, particularly the first two, is

important to the successful completion of the mission, their

use was carefully programed prior to flight to allow a suf-

ficient reserve in the event of a procedural change during

the mission.





7 - 56 G O P f F I D E f T T l A r

was recommended that at least 5 minutes be allowed for the

gyros to achieve a good operating speed prior to uncaging.

In all cases the pilot switched properly and allowed more

than 7 minutes after powering up before uncaging the gyros.

Experiments procedures and performance: The flashing light

was deployed on schedule and was within acceptable limits in

attitude. The pilot achieved a good retroattitude and then

pitched up at 0.5°/sec. The beacon was deployed at a space-

craft pitch attitude of -20° ±1°. At that time the roll

attitude was 3% the yaw attitude was 8% and attitude rates

were very near zero. The subsequent power down and yaw maneu-

ver to observe the flashing beacon were completed as planned;

however, a possible error in yaw determination during the

first night observation probably caused the pilot to miss

this sighting. Thereafter, the flashing-light observations

were completed exactly as planned.

Six of the ten planned radiation measurements were accomplished

as scheduled. The final two measurements were not attempted

because of more pressing operational problems beginning with

the early actuation of the 0.05g relay. The other two measure-

ments which were not taken as planned were overlooked by the

pilot; the first during an attempt to sight the flashing beacon,

and the second while he was encountering difficulties with the

condensate transfer system.

The ground-light observation was completed on the 6th orbital

pass generally as planned. Continuous recording during this

experiment was not initiated;, and as a result no analysis ofattitudes and rates was possible.

The balloon experiment failed for technical reasons, but the

two attempts by the pilot to position the spacecraft for de-

ploying the balloon could be readily analyzed for performance.

During deployment, the spacecraft was to be held as close to

zero rates and attitudes as possible to allow the balloon

to trail directly rearward along the trajectory. On the first

attempt to deploy the balloon, attitudes and rates were es-

sentially zero. During the second attempt, rates were zero

with attitudes of 8Qin pitch, 5° in yaw, and 0° in roll. In

both instances, all switch functions were correct and con-ducted in the proper sequence, including the setting up of

the l6-mm movie camera.

The pilot was very conscientious in correctly performing the

dim-light-phenomenon experiment. Maneuvering for this experi-

ment consisted of alining yaw on the setting sun while holding

pitch at -3 -% caging the gyros, and going to the gyro-free

position. A roll maneuver of 3^° was then required to aline



C O N F I D E N T I A L P a g e 7 - 57

the pitch plane of the spacecraft with the plane of the eclipticfor the first series of zodiacal-light photographs. The space-craft was held by the ASCS in this special alinement by firstcaging the gyros and then switching to the gyro-free positionwith pitch torquing on. At the conclusion of the zodiacal-lightphotographs, the gyros were switched to the "slave" position,

which allowed the horizon scanners to return the space-craft slowly to its normal orbital attitude.

The pilot accomplished all maneuvers and switching as planned.Using FBW-low, he yawed over to the setting sun; however, hewas not satisfied with his alinement and introduced a correction.The 3^-° roll maneuver was difficult to monitor on the instru-ments because of the bright sunlight streaming directly intothe window. After uncaging, the roll angle was set in correctly.Since the roll maneuver had introduced a yaw component, thepilot adjusted yaw attitude before the final gyro caging anduncaging and then went to ASCS control for holding on the

ecliptic plane. After completing the zodiacal-light photo-graphs, the pilot switched the gyros to "slave" and correctedthe spacecraft back to the normal orbital attitude. For bestphotographic results, the cockpit was completely darkened;thus, it was necessary to count to himself to obtain the correctexposure times and intervals between photographs. He performedthis entire experiment in an exceptionally efficient manner.

Shortly after the dim-light photography, the pilot executeda series of yaw maneuvers to accomplish the horizon-definitionphotography. The pilot used the gyro indicators selectivelyto locate precisely the 90° positions in yaw. The maneuvers

were completed as planned and required 8 minutes during whichtime only 0.255 pound of control fuel was used.

The U. S. Weather Bureau infrared photography was performed

during drifting flight, with small manual proportional control

inputs occasionally applied to help in alining on the more

interesting cloud features. The pilot's very conservative

fuel usage and low-order control inputs in utilizing this

control mode indicated a very refined manual technique and

skill.

In spite of the 0.05g-relay problem, the pilot completed one

HF antenna test with good results. The pilot also photographed

interesting terrain features, and again the technique and pro-

cedures used reflect quite favorably on the pilot's performance.




Page 7 - 58 ~ COftTFID

7.2. .5 Pilot performance summary: The pilot's inflight performance

can be summarized as a conscientious adherence to the planned

tasks, as well as the application of proper corrective actions

when system or operational problems occurred. The pilot was

in complete control of every situation throughout the flight,

and he exhibited smooth piloting abilities. He completed his

inflight activities almost exactly as they were scheduled,

even though adjustment of the suit environmental circuit plagued

him constantly throughout the flight. The program of scientific

experiments, until pre-empted by system difficulties, was con-

ducted very satisfactorily. The premature 0.05g function

engrossed the pilot with critical operational considerations

in order to complete the flight successfully, and thesecon-

siderations took precedence over nonoperational activities.

Thereafter, the pilot elected to perform only those flight-

plan tasks which would not complicate or interfere with

management of the malfunctioning systems. Switching was both

'timely and accurate throughout the flight. Emergency switchingafter the 0.05g telelite illumination and ASCS inverter problems

became evident was excellent. Some switching in noncritical

areas was overlooked occasionally, but usually because the

pilot was concentrating on more important tasks. Both his

technique and fuel usage during manual control were commend-

able. The pilot's use of the'manual proportional control mode

was so excellent as to warrant a reappraisal of the value of

this system. Control systems checks were accomplished quickly,

in less than 1 minute per system on the average, and each check

required only approximately 0.02 pound of fuel. The pilot con-

sistently used less than the expected amounts of consumables;

and, in the absence of the system malfunctions which were ex-perienced, he would, as planned, have had more than enough con-

sumable quantities to accomplish easily all of the many flight

activities.

C O N r i D D M T I A f c



JVF1DEIVTIAL"

-

Page 7-59

TABLE 7.2.1.1-1.- PILOT TIME IN SPACECRAFT 20 DURING

HANGAR AND LAUNCH COMPLEX TESTS

Date

Oct. 11 to

19, 1962

Nov. 11, 1962

Jan. 5, 1963

Jan. 12 and

Mar. 1, 1963

Mar. k, Ik, 15,1963

Mar. 19, 1963

Mar. 20, 21, 22,

1963

April U, 1963

April 18, 1963

April 23, 1963

April 2 , 1963

May 3, 1963

May 6, 1963

May 8, 1963

May 10, 1963

May 14, 1963

Total

Test description

Integrated systems tests

RCS - hangar

Altitude chamber

TV systems test

Communications systems radiation test

Darkness and egress

Simulated flight, hangar

Prepad RCS test

Alinement, weight, and balance

Systems test and simulated flight no. 1

Electrical mate

Mark instrument normal and emergency limits

Flight configuration sequence and abort

Launch simulation and RF compatibility

Systems test and simulated flight no. 3

Countdovn (canceled)

Duration,hr:min

06:45

03=15

06: 5

07:00

04: 5

01:20

12:10

00:50

0 :00

0 :00

0 :30

00: 5

03:00

05:00

03: 5

06:00

73:50




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Page 7-62

TABLE 7.2.1.5-1.- PILOT PKEFLIGHT ACTIVITIES

FROM JANUARY 1, 1963 TO LAUNCH DATE

Date

Jan. 2

Jan. U to 7

Jan. 10

Jan. 12

Jan. 18 and 19

Jan. 21

Jan. 22

Jan. 23

Jan. 2

Jan. 25

Jan. 30

Jan. 31

Feb. 1

Feb. 2

Feb. 3

Feb. k

Feb. 5

Feb. 6

Feb. 7Feb. 8

Feb. 11

Feb. 12

Feb. 20

Feb. 21

Feb. 23

Mar. 1

Mar. U

Mar. 6

Mar. 8

Mar. 12

Mar. 13

Mar. lU

Mar. 15

Mar. 19

Day

Wed.

Fri. to Tues.

Thurs.Sat.Fri. and Sat.Mon.

Tues.

Wed.

Thurs.

Fri.Wed.

Thurs.Fri.Sat.Sun.Mon.Tues.Wed.

Thurs.

Fri.Mon.

Tues.

Wed.Thurs.Sat.

Fri.Mon.

Wed.

Fri.Tues.

Wed.Thurs.

Fri.

Tues.

Q

Activities

Altitude Chamber Systems Test Review,

blood-pressure checkout in altitude

chamber, flying (TF-102A)

Altitude Chamber Systems Test

Flight-plan review, flying (TF-102A)

TV systems test, flying (TF-102A)

Morehead Planetarium (celestial reviev)Weight and balance

Systems briefings (ASCS andRCS)

Systems briefings (communications andsequential)

Flight- plan and experiments reviewSystems briefings (electrical andECS)

Flying (F-102A)

Flying (T-J5A)

Booster rolloutFlying (T-JJA)Flying (T-3JA)

Experiments status review

Flight -plan reviewCouch fitting

Flying (T-JJA)Observation of flashing beacon on T-33A

Flight-plan briefing to Deputy Directorfor Mission Requirements

Flying (F-102A)

Flying (F-102A), flight-food testingExperiments briefingsFlying (T-33A)

TV systems test

Communication systems radiation testWeight and balance

Flying (F-102A)

Couch fitting

Flying (T-JJA, F-102A)

Communication systems radiation testCommunication systems radiation test,Mercury Procedures Trainer

Darkness and egress test

Includes only major activities and does not include such activitiesas spacecraft reviews, physical exercise, study, monitoring systems test,

informal briefings with operational and systems personnel.



I E N T I A L Page 7 - 6 3

TABLE 7.2.1.5-1.- PILOT PREFLIGHT ACTIVITIES FROM

JANUARY 1, 1963 TO LAUNCH DATE - Continued

Date

Mar. 20 to 2k

Mar. 2

Mar. 26

Mar. 27

Mar. 28

Mar. 29

Apr. 1 and 2

Apr. U

Apr. 5

Apr. 6

Apr. 7

Apr. 9

Apr. 10

Apr. 11

Apr. 15

Apr. 16

Apr. 17

Apr. 18

Apr. 19

Apr. 22

Apr. 23

Apr. 2}4

Apr. 25

Apr. 27

Apr. 29

Apr. 30

May 1

May 2

May 3

May ^May 5'4ay 6

DayWed. to Sun.

Sun.

Tues.

Wed.

Thurs.

Fri.Mon. and Tues.Thurs.Fri.

Sat.Sun.

Tues.

Wed.

Thurs.

Mon.

Tues.

Wed.

Thurs.

Fri.

Mon.

Tues.Wed.

Thurs.

Sat.

Mon.

Tues.

Wed.

Thurs.

Fri.

Sat.

Sun.

ton.

Activities

Simulated flight (Hangar)

Flying (F-102A)

Flying (T-33A)

Flying (T-33-M, Mercury Procedures Trainer

Flying (T-33-M, Centrifuge - acceleration

re familiarization

Mercury Procedures TrainerMercury Procedures TrainerDOD-NASA MA-9 Review, Prepad RCS test

Mercury Procedures Trainer, flying(TF-102A), Morehead Planetarium(Celestial review)

Morehead Planetarium (Celestial review)Flying (F-102A)

Flying (F-102A)

Egress and recovery training

Egress and recovery training, survival

pack exercise

Flying (F-102A)

Mercury Procedures Trainer, mission and

flight controller briefing

Mission and flight controller briefing

Alinement, weight, and balance, MercuryProcedures Trainer

Mercury Procedures Trainer

Mechanical mate

Simulated flight no. 1

Electrical mate

Mercury Procedures TrainerMercury Procedures Trainer

Yaw demonstration (AF Hangar)

Systems briefings (review)

Systems and operations examination

Launch simulation, Mission Rules review

Examination questionnaire review, markedspacecraft's normal and emergency

instrument limits

Launch simulation

Flying (TF-102A)

Flight configuration sequence and aborts

Includes only major activities and does not include such activities as

spacecraft reviews, physical exercise, study, monitoring systems test, inform

briefings with operational and systems personnel.








Page 7-66 IPCMTM.LI

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T A B L E 7.2.2.2-1.- PILOT'S E Q U I P M E N T LIST

Special equipment storage kit

Hand-held Hasselblad camera

Hand-held Ro~bot camera

U. S. Weather Bureau film magazine

MTT film magazine

Six-inch focal length TV lens

O.JO neutral density filter for 6-inch TV lens

Special water flask

Water-tight container

Condensate purification containers

Vomitus container

Food

Writing desk

l6-mm balloon film magazine

l6-mm reentry film magazine

Extinction photometer

Standard calibration light

f/2.8 50-mm lens for l6-mm camera

Exposure meter

Sighting device for the Hasselblad camera

Food

Star navigation deviceNavigation booklet

Glove box

Flight-plan roller and flight plan (see fig. 7-2.2.2-3)

U . S. Weather Bureau filter mosaic

Radiation dosimeter

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WriDCNTIAL Page 7-73

TABLE 7.2.4.1-1.- CONTROL MODE USAGE

Control mode

configuration

Drift

Drift and MP

ASCS orbit

Drift and FBW-low

FBW-low (gyros uncaged)

ASCS reentry

MP (gyros uncaged)

Percentagetime used

in rank order

43

26

13

13

2

2

1

Maximum time

used at any

one time,

hrrmin

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8:44

1:20

3:11

0:11

0:37

0:04

Frequency

used

2

1

7

l

8

1

5

Note: 1. Percentage usage times are test estimates "because

the tape recorder was in the program mode during

some switching operations and because of the

apparent tape failure during the rest period.

2. There is only intermittent use of the control

system during all of drift and MP and drift

and FBW-low periods.

C O N F I D E N T I A L•



7 7 ^

7-3 Pilot's Flight Report

7.3.1 Introduction.- The MA-9 mission was designed to evaluate

the operational and aeromedical aspects of the man-

spacecraft system in a manned one-day mission. The flight

of Faith Seven has demonstrated this capability and, I

hope, has shown the way for expanded capabilities in the

manned space-flight program.

Because of the extension of the flight duration for this

flight, there was much more time available for evaluating

systems, conducting experiments, and determining man's

adaptability to the space environment.

The name of my spacecraft, Faith Seven, was chosen for

three reasons; a belief in God and country, a loyalty tomy parent organizations, and a confidence in the entire

Mercury space team.

This flight was an excellent example of a well coordinated,

well disciplined, voice procedure between the spacecraft

and the world-wide tracking network. The information

received onboard Faith Seven was clear, concise, and served

to keep me continuously updated throughout the flight.

In my opinion, it is pertinent that this data flow was so

complete that I felt confident of my ability to accomplish

the mission successfully in the event of an unexpected

loss of air-ground communications.

The remaining context of this report will be a chronological

discussion of the flight from launch to recovery.

7.3.2 Launch.- Launch activities were well planned and were accom-

plished smoothly. We were always on or slightly ahead of

schedule, and I did not feel at all rushed for time. The

donning of the Mercury pressure suit for my second flight

attempt was accomplished in good order, and we arrived at

Launch Pad Ik on schedule. I was inserted into the

cockpit, and we had completed the necessary checks

7 minutes ahead of the scheduled time for gantry removal.This time, the gantry came back without any delays, and

the countdown proceeded very successfully with lift-off

occurring 4 minutes behind schedule.

The time interval between my insertion into the spacecraft

and lift-off did not seem excessive, although insertion

GONriDCHTIAL





Page 7-76

would; however, the thrust of the posigrades gave me a

distinct boot in the rear.

Window discoloration, or smudging, occurring during powered

flight. The window appeared to be smudged in two areas.

The first smudge was streaks, similar to powder burns, on

the outside of the window. The second appeared to be a

solid, greasy, coating on the inside of the outside pane

of glass. It appeared oily when I looked through it with

direct light; but with oblique light, it blotted out the

external view and resembled ice or frost.

7 - 3 - ^ - Orbital flight.- The turnaround maneuver felt just as it

did on the procedures trainer. I went to auxiliary damping

control mode, waited for a short interval, and then selected

fly-by-wire control and started a k°/sec left yaw rate using

instruments. I did not pitch down because we had calculated

that a left yaw rate would result in a slow negative pitch

rate. However, the spacecraft did not pitch down as much

as expected during the turnaround; and, consequently, my

pitch attitude was still well above orbit attitude at com-

pletion of the maneuver.

Immediately after turnaround, my attention was attracted

to the booster, which was not more than 200 yards away.

I could read the lettering on the sides and could see

various details of the sustainer, such as the tanks. It

was a very bright silver in color, with a frosty white band

around the center portion of it. It was still wisping

lox and fuel from the aft end. It was yawed approximately

15° to 20° to its left. I had it in sight for a total of

approximately 8 minutes. The front end was slowly turning

in counterclockwise rotation'. The last time I saw it, it

was turned about 70° to the azimuth at which I had departed

from it. I began to drop slowly down to my -3 -°pitch

attitude. After selecting ASCS control, I pulled the

l6-mm camera out of its bracket and took a short film

burst of the booster with part of the east coast of the

United States and Cape Canaveral in the background;

however, the booster had moved some distance away by thistime.

I agree with Scott Carpenter that visual perspective

changes when you go into zero-g. The cockpit did seem to

be somewhat differently located in perspective to myself.

You move up forward in the seat, regardless of how tight

your straps are cinched. The ditty bag (special equipment

storage kit) on the right seems to be a different angle

to you than it is when you are on the launch pad. I did



WriDCNTIAL Page 7 - 77

feel very distinctly that I was sitting upright. Most of

the time I felt as if I were lightly floating. A couple

of times I felt almost as if I were hanging upside down

because of the feeling of floating into the shoulder

straps. Every time I "dropped" something, I grabbed below

it, expecting it to fall, (items stayed close to the

place I put them when the cabin fan was off.)

Speed is very apparent when overflying clear or broken-

cloud areas. You definitely have a feeling of really

traveling along. If the cloud is a solid deck underneath

you and you don't have any other motion cues, you have a

very slow, floating feeling.

You really need to use automatic control on the first

orbital pass or to have a low work load in order to

collect your senses, to acclimatize yourself to this new

situation, and to organize the flight activities. I felt

that I was not on top of the situation as completely as

I would like to be right after insertion. Although I was

thinking about all the items to be done and of how to do

them, I did not feel completely at home. I felt that I

was in a strange environment and was not at my best, until

perhaps halfway through the pass. By the end of the first

pass, I was feeling really ready to power down and go into

drifting flight - able to manage the spacecraft in manner

or means.

The operating bands of the automatic control mode, particu-

larly in yaw and roll, appeared to "be wider than I thoughtthey would be. You just slop along between these 11-5° bands,

Sometimes, it varied right out to the full extent of the

limits. It does not hold a precise attitude, as I knew

it would not from a previous engineering analysis. However,

it did not occur to me that these excursions would show

up so much from visual reference. You soon get accustomed

to it, and it is no problem.

I checked my T +5 second relay by going to gyros-frees

with pitch-torquing on, and it worked. I got TV on for

the Canary Islands, gave a consumable readout over Zanzibar,

and checked the manual proportional system and found that

it was functional. The manual proportional control mode

in procedures trainer felt identical to that in the space-

craft. It worked very well. After becoming used to the

slow light-off and the slight lag, I could control attitudes

easily, even down to small changes. It is not as good

for very fine attitude control as the 1-pound fly-by-wire

low thrusters.

•



Page 7-78 1»• J.J^fHill I

I put a short status report on the tape at ho minutes

and went into the night side. I immediately saw the

white haze layer, as described by Wally, in which the

stars would fade out, when behind it, and then reappear

below it before disappearing behind the horizon. The

earth has a sharp horizon even at night. I estimated the

haze layer to be about 6° or 7° in height and located

2°, or perhaps 5° above the horizon line.

I never tired of looking at the sunsets. As the sun

begins to get down towards the horizon it is very well

defined, quite difficult to look at, and not diffused as

when you look at it through the atmosphere. It is a very

bright white; in fact, it is almost the bluish white color

of an arc. As it begins to impinge on the horizon line,

it undergoes a spreading, or flattening effect. The sky

begins to get quite dark and gives the impression of deep

blackness. This light spreading out from the sun is a

bright orange color which moves out under a narrow band

of bright blue that is always visible throughout the day-

light period. As the sun begins to go down, it is re-

placed by this bright gold-orange band which extends out for

some distance on either side, defining the horizon even

more clearly. The sun goes below the horizon rapidly, and

this orange band still persists but gets considerably fainter

as the black sky bounded by dark blue bands follows it on

down. You do see a glow after the sun has set, although

it is not ray-like. I could still tell exactly where the

sun had set a number of seconds afterward.

I conducted the emergency voice check over Muchea, sent a

blood pressure, switched the S-band beacon to "ground

command," and at 1:10:00 gave a short status report.

When there is no moon, the earth is darker than the sky;

there is a difference in the two blacks. In general, there

was more light from the sky; the sky is a shining black

as compared with a dull black appearance of the earth.

There is a distinct line at the horizon and the earth is

the darker.

I saw the lights of Perth on the'west coast of Australia.

If there is moonlight, then cloud layers and ground

features can be seen. The moonlight is bright enough to

see the motion over the ground. On several occasions I

could see light from cities on the ground glowing through

the clouds.

•-*-3 ' "

CONFIDENTIAL





Page T - 80

good yav indications. At night I saw some towns andcities underneath some of the cloud decks, which areexcellent yaw indicators.

At 3:25:00 I went to fly-by-wire low, slowly pitched up

to the -20° mark on the window, deployed the flashingbeacon, and there was a loud "cloomp" as the squib firedand it departed. I then caged the gyros and powered downthe ASCS a-c bus. I never did see the beacon on that firstnight, but I was having some difficulty finding my180° point. I tried unsuccessfully to observe the flash-ing beacon early on the day side also.

On the second night side after deploying the flashing beacon,shortly after going into the night side, I spotted thelittle rascal. It was quite visible and appeared to beonly 8 to 10 miles away. I deliberately moved off target,

waited until 5: 0:00 and eased back to l80° yaw andsaw the light again, at which time it appeared to bearound 12 to 1^ miles away and still quite visible.

On the third night side after deploying the flashing light,I had, no anticipation of seeing it at all; but at 6:56:00

there it was, blinking away. It was very faint andappeared to be at a distance of about 16 to 17 miles. Iwould say it was approximately the brightness of a fifth-magnitude star, whereas on the second night side afterdeployment it appeared to be about that of a second-magnitude star.

At 4:25:00 I gave the medics their first orbital urinesample. At 4:5 :00 I ate four brownies from one of thelittle snack boxes and drank five or six gulps of water.While in flight, you must force yourself to eat and drink.When you have a fairly full flight plan, the temptationis not to bother with eating and drinking, but to devoteyour time to doing a good job on the items you have to doso that you can do them correctly. I deliberately madea point of forcing myself to drink water regularly. Thefood was so difficult to prepare and to eat that I did noteat the quantity of food I had planned to. The older type

of tube food would have been better on this flight.

Control of suit circuit is definitely marginal. It wasphysiologically and psychologically the worst problem Ihad during the flight. I was concerned with the suit circuit,probably more than with any other system throughout theentire flight. I worked continuously to keep it withinlimits. The suit was very moist and I was really soakedwith water for some time before recovery. Controlling the



nP a g e 7-81

suit heat exchanger was a big problem during the entire

flight. Wo setting, regardless of how small the change,

would hold the suit dome temperature within acceptable

limits. It was frequently frozen, at which point I would

have to turn the suit water flow completely off until

it thawed. At times, it would then go to the high temperature

side, and the suit-inlet temperature would also become

uncomfortably high. I would slowly work it back down to a

point where I was beginning to get a comfortable suit; and

then upon leaving it there, it would hold for a short while

and suddenly would plunge on down to the freezing mark and

I would have to start all over. Even when I would remember

the settings that I had previously used and go back to

even slightly lesser settings than these, it would hold

for a short while, but then plunge or rise. It did not

appear to be constant at all. Naturally, opening and

closing the visor also added to the variations.

At 8:21:00, I pitched down to observe the ground light with

gyros free. At 9:00:00, I tried to deploy the balloon,

with no success. At 9'10:00; I tried again to deploy the

balloon, and again with no success. I used the procedure

on the checklist, but nothing happened.

The valve on the drinking container leaked so much water

that I could not place water in any of the plastic food

containers. The plastic food containers, with the frozen

dehydrated food, are completely unsatisfactory for zero-g

use. Under zero-g, there is no way of getting the plastic

container away from the nozzle to work the water down into

the food. The water tends to come out of the plastic

top as you try closing it off. I tried one of them and

had so much trouble and got water all over myself, my

gloves, and the instrument panel, but only enough water

in the container to wet approximately one-third of the food

that was in there. I finally just gave it up as a lost

cause, and ate only the snack-type foods.

The condensate bag appeared to fill up much sooner than it

should have and was so hard to pump against that I was afraid

I would rupture it if I pumped any harder. I finally

stopped pumping, transferred over to the 4-pound tank andproceeded to pump some water into it. Subsequently, the

pump appeared to be jammed, and after switching back to

the other tank the needle fitting came out with the entire

system completely out of commission.

As previously stated, I found that orienting the space-

craft after drifting flight was quite easy on the day side

and not too difficult on the night side, although





i O M r i D C N T l A L P a g e 7 - 83

I dropped off to sleep very soundly at about 13:30:00, andslept on until I woke up at 1 :48:00. I dropped off tosleep again at about 16:50:00 and slept right on throughuntil 17:50:00. I woke up and the dome temperature light

was on. After adjusting the temperature, I dropped offto sleep immediately thereafter and woke up at 18:25:00.

I dropped off to sleep again and did not wake up until19:27:00.

During portions of the rest period when I was awake, I madeseveral interesting ground observations. I could detect in-dividual houses and streets. I also saw what appeared to betrains and trucks in some of the clear regions in the Tibetianarea. I noted several cases of observing wind direction and

velocity on the ground from smoke emanating from smokestacksor fireplaces of houses. Particularly in the Himalayaarea I could see houses, yards, fields, roads, streams, and

lakes. I could see a lot of snow on the ground in the upperportions of the mountains and a lot of the lakes frozen overeven down in the lower sections of the windblown, sandy, highplateau areas of the Himalayas.

During the day, the earth has a predominantly bluish cast.I found that green showed up very little. Water lookedvery blue, and heavy forest areas looked blue-green. Theonly really distinctive green showed up in the high

Tibetian area. Some of the high lakes were a bright emeraldgreen and looked like those found in a copper-sulphatemining area. The browns of the Arabian desert showed up

quite distinctly, but the Sahara was not quite so brown.If you are looking straight down on things, the color is

truer than if you're looking at an angle.

At 21:34:00 on the night side, I observed a line of lightedcities along the East Coast of Australia - Sidney, Melbourne,Brisbane. I powered up the ASCS bus at 23:30:00, alinedthe spacecraft at 23: 0:00 and uncaged the gyros. The Cap

Com in the Mercury Control Center mentioned that the scanneroutputs and gyro outputs agreed perfectly.

At 2 :15:00, I went to fly-by-wire low and zeroed the space-

craft yaw axis on the setting sun,which was very difficultbecause the sun is extremely bright. I then caged thegyros, brought them back to free, rolled to 3^° right,caged the gyros again, brought them back to free. Thepitch torquing switch was on and I placed the spacecrafton automatic mode to start the zodiacal light photographs.




Page 7-84

On the latter part of the night side, I placed the gyrosto "slave" and the scanners began to correct the spacecraftback to the orbital plane. Meantime, I was busily snappingpictures. The fuel quantity warning light came on at6l percent at 24:58:00. I remained on automatic control,

snapped two pictures of the horizon-definition quadrantphotographs, went to gyros free, yawed 90° right, tooktwo pictures, caged and uncaged the gyros, went to 90° rightagain (which put me directly into the sun)snapped twopictures, went to 180° right (which put me to the270

0point) snapped two pictures., gave one negative yaw

pulse to start back around to get the gyros off thel80° stop, let it drift on around for a few moments, cagedthe gyros, and powered down.

At 25:20:00 the moon was beginning to set in the west. Iwas faced in the right direction so I used manual propor-

tional to keep my attitudes nearly correct and made threeshots of the setting moon for M.I.T. At this time themoon appeared full from earthshine.

At two different times, I saw a faint glow just aftersunset or prior to sunrise; it was somewhat cone shaped,and I believe it was the faint glow of zodiacal light. Itwas not exactly vertical to the horizon. I had a feelingthat this was just a glow off the sun. It was not asbright as the Milky Way. Another night phenomenon that Inoticed occurred when I was over South America looking Eastor Northeast. At the particular time I couldn't see this

layer, but I had the feeling that it was more of a ceilingthan a layer. It was not distinct and did not last long,but it was higher than I was, was not well defined, andwas not in the vicinity of the horizon. It was a goodsized area very indistinct in shape. It had a faint glowwith a reddish brown cast. It seemed to be quite extensive,very faint, and contrasted as a lighter area in thenight sky.

On the 17th and l8th orbital passes, I took infraredphotographs. The G.m.t. time check over Mercury Controlat 26:45:00 showed the G.m.t. clock to be 10 seconds fast.

I made more photographs. I was using full drifting flightand engaging the manual proportional handle to make somevery slight attitude corrections when necessary tophotograph, trying to hold a window-down attitude, andallowing yaw to ease around wherever it would. At timesduring the day, I did note that the sun was very, veryhot through the window. The particular pattern of thesun would be hot on my suit. I could feel heat on theinside of the window from the sun, too,and through myglove.

CONFIDENTIAL





7 - 8 6

7.3.5 Retrofire.- As a result of the premature latching in of

the ASCS 0.05g relay and the subsequent loss of the ASCS

main and standby inverter power, I was required to initiate

and control the spacecraft manually during the retrofire

and reentry events. I decided to control the spacecraft

during the retrofire period on manual proportional with

fly-by-wire ready as a backup, if necessary. I stowed the

equipment and completed the necessary preretrofire switching

procedures well in advance and was all set except for the

squib switch.

As I entered the last night side prior to the retrofire

event, and realizing that I would have to orient the

spacecraft to the proper retroattitude within 10 minutes

after sunrise for a countdown from Coastal Sentry Quebec

(CSQ), I decided to orient the spacecraft to retroattitude

during the night side and stay fairly close to it. I was

able to use stars that were near my flight path as well

as cloud patterns after the moon came up to maintain proper

yaw orientation. Shortly after sunrise when sunlight

first struck the window at an oblique angle, the window

became completely translucent, thus making it impossible

to see through it for aljnost a minute. I flew instruments

(rate needles) during this period and when I again was able

to see out the window I was still close to the proper

retroattitude. I positioned the spacecraft exactly on

retroattitude. I could now see out the window clearly,

and the manual proportional control system was operating

very well.

The procedure from the ground to prepare me for retrofire

was good with one exception. The original plan had been

to count down to retrosequence (if it was to be used) and

then to retrofire. When retrosequence was not used, I

had rejected the idea of any count except to retrofire,

so that there was some doubt in my mind when the CSQ

Cap Com began a countdown, whether it was retrofire or not.

I would recommend that we not count down in this same manner

for anything but retrofire in the future. Fortunately,

we had set the clock on minutes and seconds so that I

assured myself with this and waited for the second count.

The CSQ spacecraft communicator (John Glenn) gave me a

countdown for retrofire, and at 5 seconds to go, I placed

the squib switch to "on." I punched the retrofire button

at time zero and the retrorockets commenced firing right

on time.



Page 7 - 87

Because of the loss of my attitude indicators, I wasrequired to use the view of the earth through the windowfor attitude reference. During the thrust period of theretrorockets, I used the gnat rate gyro indicators inorder to control rates while crosschecking through thewindow for attitude control. This method proved to Toe

somwhat difficult because of the high contrast in brightnesslevel existing between the exterior and interior of the

spacecraft. Consequently, I had to hold up my hand toshade my eyes from the light coming through the windowwhile viewing the rate indicators. Other than this, Ihad no difficulty in controlling the spacecraft during thisevent.

The first retrorocket offset moment was fairly mild; Ikept the rates close to zero and observed that attitudes

were still fairly close to nominal. As the second onefired, the attitudes shifted off a little but I broughtthem back and then shifted my vision to the rate indicators.The greatest offset moment appeared to be in plus yaw whenthe number two retrorocket commenced firing.

I had plenty of training using the fly-by-wire and manualproportional control modes to control retrorocket firing

disturbances. I did not have the opportunity to practicecontrolling retrofire simulations on the ground trainers

using a combination of window reference and rate indicators.This, however, was not a great problem.

If I had lost air - ground communications prior to retro-fire, I would have been able to use the last retrofire

time I received, which was correct, and could have accomplished

retrofire at the proper time and in a proper manner.

The retros give a good, solid "thump" in the seat of your

pants and I could very easily count each one as it ignited.Retrojettison, which I did manually, was a very solid

"clack" in the retropack area. I felt that I could actuallyfeel the pack depart.

I maintained retroattitude for part of the 10-minute period

prior to the nominal 0.05g time in order to keep the earthreference in view as long as possible before going to

reentry attitude. A few minutes prior to the time for

0.05g, I slowly pitched up to a negative 12° to 1^° pitchattitude, just keeping the earth horizon at the bottom ofthe window and holding yaw and roll at zero.



Page 7-

7.3.6 Reentry.- Initially, I was intending to use fly-by-wire

for reentry. The fly-by-wire pitch-up high thruster was

slow to light off. So I pushed in the manual proportional

handle and decided to go dual authority, which gave me

9 pounds of thrust capability in the pitch and yaw axes.

About 1 minute prior to 0.05g, the spacecraft began to

feel definitely like it wanted to reenter. It was sluggish

on the controls, and began to pitch up to zero pitch

attitude. I allowed it to pitch on up, and started a

negative roll rate. I was rather surprised to see what

a "wallow" the spacecraft set up with this roll rate.

This, of course, is something you do not see in the procedures

trainer. The spacecraft more or less "wallowed" or spiraled

around, and the pitch and yaw rates began to build very

slowly; they were a very low order of magnitude. For the

early oscillations, pure manual proportional would have

handled them very nicely. The 1-pound thrusters would

almost handle the initial oscillations. Later the

oscillations had considerable more force to them, and

required more thrust to damp them.

The fly-by-wire high thrusters were not used during the

first part of the reentry. The oscillations began to

increase and thereby required a continuously greater control

input. It was at this time that the fly-by-wire high

thrusters unexpectedly fired, causing me to overshoot two

different yaw oscillations. I began to be a little more

cautious on the amount of thrust called for and had them

well under control prior to max g. On down through max g,

I held the rates down relatively low. At 95;000 feet on

down to about 50,000 feet, the rates became quite pronounced.

It seemed to take on a different ratio in amplitude and

frequency. Even using dual authority with ^9 pounds of

thrust, I still was not able to pin the rates as well as

I would have liked to. The oscillations were held to a

reasonable degree until about 50,000 feet at which time

the oscillations got fairly large and fast and I was really

having difficulty controlling them. You could actually

feel the oscillating g-forces, but they were not physio-

logically objectionable at all.

7.3.7 Landing and recovery.- As planned, I started to deploy the

drogue parachute when passing through 42,000 feet as

indicated on the altimeter. I pushed the drogue button

and the drogue came out immediately. I realized I was

in the clouds when the parachute deployed, but I still

could see it very easily. It immediately stabilized the

spacecraft and it looked very nice.



WIAL* Page 7 -

The main parachute came out automatically at approximately11,000 feet indicated altitude, reefed normally, and then

blossomed fully.

The g's from the parachute opening were much less than Ihad expected. Rate of descent after parachute opening was

between 35 and 40 feet per second. The spacecraft wasoscillating slowly under the parachute. At 4,000 feet

indicated altitude the rate of descent was down to JO feet

per second, but oscillations were still present.

The landing bag deployed automatically. Landing was solid

but not severe. Considerable water splashed in, apparently

through the snorkles, but was splashing all about the

cockpit in small amounts from each side. The spacecraft

went down to the left side, rolled around with my head

down, and wound up with the right side of the spacecraftunder the water. I could see the spacecraft then ease

back up the water line where it was lying flat in the water,with my head up.

The helicopter pilot circling the spacecraft reported that

the parachute was slow in disconnecting. I put the main

and emergency disconnect fuse switches to number 1, and

the rescue aids switch was placed to "manual." (l am

not absolutely certain, but it is my feeling that I placed

the rescue aids switch to manual immediately after landing.)It appeared to be a matter of a minute before the space-craft began to right itself, and it came right up to an

upright position.

At this point, I used the swizzle stick and turned the

manual-fill nitrogen handle, pulled it out, turned it tothe "on"position and then put the rescue aids switchto "automatic" to extend the whip antenna. By this time

there were swimmers actually in the water around me.

Immediately after main parachute deployment, I began tohear the helicopters contacting me on radio. There were

two of them circling me at this time. They stated thatthe carrier was very nearby (4.4 miles) and would be

right with me. This was extremely comforting. I was

asked how I preferred to be recovered. I elected, as I

had planned, to be hoisted onboard the carrier.

Everything proceeded very smoothly, and I did not get outof the spacecraft until John Graham and his troops blew

the hatch from the outside on the carrier. I was met byJohn and Dick Pollard, and we took two blood pressures

C Q N r i D E P f T I A L




while I was still inside the spacecraft. I then climbed

out and we took another blood pressure. While standing

still for the blood pressure I "began to feel somewhat

light headed. This feeling disappeared as soon as I had

taken two steps away from the spacecraft. At no time

after that did I again notice any more light headedness.

We then proceeded below for the medical and technical

debriefings.

C O N r i D E N T I A L





Page 7-9



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8.0 FLIGHT CONTROL AND NETWORK PERFORMANCE

8.1 Flight Control Summary

8.1.1 Prelaunch operations.- The Mercury Control Center (MCC) and

Bermuda (BDA) flight controllers were deployed for the MA-9

mission on April 30, 19^3. The remainder of the flight con-

trol teams began to deploy on May 2, and all teams had " b e e n

deployed by May 5, 196j. The MCC-BDA flight control teamswere exercised through a series of 10 launch simulations.

These simulations exercised the decision capability of the

flight controllers and the astronaut during the criticalpowered-flight phase of the mission.

Four days of network simulations were conducted in prepara-

tion for the mission. The exercises performed on May 1, 8,

and 9 involved two simulations daily. The final networksimulation was performed on May 12, 1963. At the completion

of this exercise, the network and the flight control teamshad reached a satisfactory state of readiness and were ready

to support the scheduled flight.

Because of the extended flight duration, an additional flightcontroller was assigned to the flight control teams at both

Kano, Nigeria (KNO) and Zanzibar (ZZB). The Kauai Island,Hawaii (HAW) site was supported by two complete flight con-

trol teams, because of the frequency of its contact periods.

Several new flight controllers were used, primarily at BDA,

ZZB, Rose Knot Victor command ship (RKV), and HAW. The CorpusChristi, Texas (TEX) site was again used as a training facil-

ity for the mission. Five aeromedical monitors and five

procedures-trainer personnel from the Flight Crew Operations

Division participated in the simulations at this site.

The Mercury Control Center was manned on a two-shift basis.

The shift changeover occurred approximately every 8 hours

throughout the flight. The flight control countdown was in-

itiated by the second shift team at T-215 minutes. The first

shift was on duty at T-120 minutes. Both shifts were on dutyfor the launch and reentry phases of the mission.

The flight control documentation for the MA-9 mission was sat-

isfactory. A total of 3^ Instrumentation Support Instructionswas transmitted to the network. The majority of these docu-

ments required only one revision during the prelaunch period.The only major revisions required were those to the instrument

calibration curves. These changes were necessary because of

calibration shifts and are normally transmitted after the launch-

pad tests, conducted 4 days prior to lift-off. Other minor





Page 8-3

data channel and the shift register in the range digital data

channel. The simultaneous failure of both components compli-

cated an effective failure analysis and caused the earlier

difficulties in solving the problems.

At 4:00 p.m.e.s.t. it was decided that the BDA FPS-16 radar

system would be able to support the launch and that the count-

down for the second launch attempt would be initiated the

following day.

The countdown was recycled for 24 hours, and the network count-

down was resumed at 2:00 a.m.e.s.t. on May 15, 19&3- All

primary network systems, with minor exceptions, were operable

when the countdown was initiated.

The confidence summaries transmitted by the network to verify

the site and calibrations were very good. No major discrepan-

cies were noted in the network voice communications. However,

the ZZB, Canton Island (CTN), RKV,and Coastal Sentry Quebec (CS

sites were influenced by propagation, and several repeats were

required from the stations. The ships were serviced by two

diverse radio links. The CSQ communications were routed through

New York and Honolulu. In most cases, these different paths

allowed communications to the ships during the periods ofday-

night frequency transition. The sites suffered short-duration

dropouts on the voice links throughout the test; however,

communications were quickly reestablished.

This countdown was continuous, except for a short hold for the

launch-vehicle ground support equipment at T-llminutes 30 sec-onds. The countdown was resumed within approximately 4 minutes

and lift-off occurred at 8:04:13 a.m.e.s.t.

8.1.3 Powered flight.- The powered flight phase was normal and all

launch events occurred at nearly the expected time. The guidanc

and data systems performance was excellent. All data sources

provided good, consistent data. Sustainer engine cut-off occurr

at 00:05:02.9 at a space-fixed velocity of 25,735 ft/sec. The

flight-path angle at cut-off was +0.00321°. A "go"condition

was evident at insertion, and the orbit lifetime was notcon-

sidered to be a problem. All spacecraft and launch-vehicle

systems performed satisfactorily during the launch phase, andthe air-ground communications were somewhat better than those

in previous missions.

8.1.4 Orbital flight.- After spacecraft separation, the astronaut

performed a fly-by-wire (ZBW) turnaround maneuver. Shortly

after completion of this maneuver, the BDA capsule communica-

tor (Cap Com)advised the MCC that an approximate 6° rise in

C O N F I D C H T I A L " -



Page 8 -

• -»**»*- " - . • • - - • . . . . tttfg.f.,

CONFIDENTIAL

c a ~ b i n - and suit-dome heat-exchanger temperatures had been

observed on telemetry at BDA. The astronaut was appraisedof this situation and increased the coolant flow. At space-craft loss of signal with BDA, the dome temperatures hadstabilized. When the astronaut acquired communications with

the Canary Islands (CYl), he said that the dome-temperaturewarning light was on. This warning indication was causedby the suit dome temperature decreasing below 51° F. Theastronaut was required to monitor this temperature through-out the flight and make frequent adjustments to the coolantcontrol valve. As a result of the exit heat pulse, thecabin temperature indication on telemetry rose from 9 -° Fat launch to approximately 118° F when the spacecraft was overMuchea, Australia (MUG). Subsequently, this temperature de-creased slowly to a value between 90° F and 100° F. Thecabin-air temperature appeared to vary slightly as a functionof the spacecraft a-c power usage. During the periods in

which the automatic stabilization and control system (ASCS)115 v a-c inverter was being used, the temperature would riseto a value of between 97° F and 105° F; and when this inverterwas not being used, the temperature would apparently decreaseslowly over a period of two passes to a value of between90° F and 97° F. All spacecraft systems were functioningnormally, and the MCC advised the Guaymas, Mexico (GYM) siteto transmit a "go" decision for seven orbital passes to theastronaut.

The first minor discrepancy occurred over MCC at the beginningof the second orbital pass. When the telemetry was commanded

by the ground, a series of repetitive R- and Z-calibrationsoccurred. It was decided that the programed R and Z calibra-tion function would be turned off during the sleep period.This anamoly occurred again when the C-band beacon wascommanded "on" over MCC at the beginning of the l6th orbitalpass.

The flashing beacon was deployed at 03: 5:00. The astronautreported that he felt the beacon deploy; however, he did notsee the light during its first night period. The astronaut

All instrumentation values presented in this section werederived in real time over telemetry, unless otherwise stated.



Page 8-5

verified that the attitudes and time were correct for deploy-ment of the beacon. The "beacon was observed during sunsetof the 4th orbital pass~"and was tracked during the night per-iod of that pass. All systems continued to perform satis-factorily. However, the astronaut was required to makefrequent changes to the suit coolant control valves in order

to maintain the suit heat-exchanger dome temperature withinlimits.

At the beginning of the 5th orbital pass, the astronautturned the cabin fan and cabin heat exchanger off as scheduledby the flight plan. It was noted subsequently that turningoff the cabin fan while in a powered-down condition did notmaterially affect the cabin temperature. The astronaut openedthe outlet port of the condensate trap in accordance with hisflight plan.

It was noted early in the flight that the actual power usage

was slightly less than predicted. This condition resulted inallowing for more radar beacon tracking during the later phasesof the flight. The C-band beacon was turned on three timesprior to passes over the HAW station to enable tracking by theRange Tracker ship. Attitude-control-fuel usage was also lessthan expected, and all reports indicated that the astronautwas managing his fuel supplies exceptionally well.

The astronaut made two attempts to deploy the balloon for theballoon drag and visibility experiment, starting at 09:00:00.All attempts were unsuccessful; and based upon analysis ofthe system and the undesirability of powering up the squibbus, it was decided that no further attempts would be made todeploy the balloon. The astronaut was advised not to actuatethe jettison switch.

During the 7th pass at 10:02:00, MCC directed ZZB to transmitto the astronaut a "go" decision to continue for 17 orbitalpasses. At this time all systems were performing well. Ingeneral, the usage of consumables was better than expectedand the astronaut was in excellent condition. Because of theexcellent performance of the astronaut and the spacecraftsystems, the task of flight control was one of monitoring

systems performance, gathering and transmitting summary infor-mation, and assisting the astronaut in carrying out his flightplan.

At the end of the 7~th pass the astronaut powered down the ASCSbus over the CSQ and began a period of extended drifting flightAt this time he stated that he was perspiring lightly and that



n fPage 8-6

he was continuing to monitor the suit dome temperature.Occasionally, the ground stations became concerned becauseof the gradual temperature increase apparent in the cabin-

heat-exchanger dome temperature. This measurement wasexpected to reach ambient eventually, but it stabilized at

approximately 75° F. The HAW site obtained good trackingon the C-band beacon on the 7th pass, and these datacon-firmed the nominal orbit. All scheduled radar tracking duringthe orbital phase was good and the acquisition and retrofiretimes were near nominal throughout the entire flight.

The RKV received a readout of the major onboard parametersfrom the astronaut at 12:28:00, and the astronaut turned thetelemetry to continuous in preparation for the sleep period.The astronaut performed a checkout of the manual and fly-by-wire (FEW) thrusters. He reported to the CSQ at 13:30:00that the system was performing satisfactorily and that he hadjust been sitting quietly looking out the window. He thenindicated that he was ready to sleep. The network sites wereadvised that the sleep period had started, and that they wereto maintain air-ground communications silence unless telemetrysignals indicated that the astronaut should be contacted.Throughout the sleep period, all parameters indicated normalsystems operation. The suit dome temperature reading receivedat the KNO site changed abruptly. The change indicated thatthe astronaut had adjusted his suit coolant flow in order tomaintain the dome temperature within the limits. The aero-medical parameters which were monitored by the remote sites

indicated that the astronaut was sleeping. The respirationrate was holding fairly steadily at 15 breaths per minute,and the heart rate was steady, at approximately 60 beats perminute. During the llth pass, the MCC received electrocardio-gram (EGG) data, remoted from the Ascension Island; and asudden increase in the heart rate from 60 beats per minute to110 beats per minute was noted. The heart rate then decreasedto 70 beats per minute. These data were interpreted by theaeromedical monitors to indicate that the astronaut may havebeen dreaming. During the llth through l4th passes, theindicated automatic fuel quantity decreased from 82 percentto 75 percent. This change initially caused some concern,

but, after evaluation it was attributed to temperature stabi-lization effects, the normal nitrogen bottle leakage rate,and the readout accuracy of the ground telemetry for automaticfuel quantity.

During the l^th pass at 21:23:00, the MUG site received a re-port on consumables from the astronaut. The astronaut stated

C O M r i D E H T l A L





Page 8 -

not yet completed, it was generally thought that the gyros,

attitude indicators,, and scanners were inoperative and that the

0.05g circuit was latched up. At this time it was realized

that a manual retrofire would be required and that a check-

list would have to be prepared for the astronaut.

At approximately 30:00:00 the remote site flight control

personnel that were on standby status were recalled to their

stations, and they were advised to stand by to attempt to

relay communications to the astronaut if directed by MCC.

While he was in contact with the CSQ, the astronaut was

requested to turn on the telemetry and the C-band beacon in

order that the Range Tracker could make a check of its radar

data, since this ship was the prime station for reentry

tracking. These data were even more important since the

retrofire maneuver would be performed manually. During the

time interval between loss of signal with MCC on the 20th pass

and acquisition of communications with HAW on this pass, addi-

tional tests were devised to verify where the ASCS logic was

latched up. As the spacecraft passed over HAW, the astronaut

was requested to place the ASCS 0.05g and emergency 0.05g fuse

switches in the on position and to select the ASCS automatic

mode to verify the 0.05g event. If the spacecraft began to

roll as it would normally do if 0.05g indications were valid,

the ASCS was latched in to the reentry mode. The astronaut

verified this roll rate, and the 0.05g event was confirmed

by telemetry over the GYM site. At this point the flight

controllers knew the configuration of the ASCS logic and

the required configuration for reentry. After completion

of this test, it was determined that the ASCS would provide

proper attitude control and roll rate for reentry.

Prior to the astronaut's acquiring communications with the

CSQ on the 21st pass, a manual retrofire checklist was com-

pleted and thoroughly checked out by MCC. This checklist

was sent via TWX to ZZB, CSQ, and HAW, and its receipt was

acknowledged. The checklist was as follows:

A. Primary procedure for retrofire

1. Attitude permission by-pass

2. Retrorocket arm switch, manual

3. Fly-by-wire thrust select switch, high and low

4. Retrosequence fuse switch, number 2



Page 8-9

5- Retrofire manual fuse switch, number 2

6. ASCS a-c bus switch, on

T. ASCS 0.05g fuse switch, number 1

8. ASCS control switch, select

9 - Mode select switch, off

10. Manual handle, push on

11. Squib arm at retrofire minus 5 seconds

12. Countdown to retrofire

IJ. Depress fire retro override at retrofire time

B. Backup to be used if there is no retrofire

1. Retro delay to instant

2. Depress retrosequence button

C. Additional precautions

1. Retropackage jettison will have to be manual.

Be sure not to arm the retropackage jettison

switch until after the rockets are fired.

2. Astronaut probably will not get a fire retro

telelite

D. Hold retroattitude and jettison retro, keep ratesas low as possible while maintaining usual referenceas an aid for low rates and at the nominal 0.05g time

(3 :09:19); select reentry mode.

E. At correct 0.05g time, select automatic on the ASCScontinuous switch.

In addition to the checklist which was relayed to the space-craft and written down by the astronaut, the CSQ attemptedto reset the clock by a ground command. The desired clocksetting for the time of retrofire was 3^ hours 59 minutes and52 seconds.

The desired time of retrofire was 33:59:30. The differencebetween the clock setting and the desired time of retrofire



- 10

resulted from a correction for the predicted clock error

which would be 22 seconds fast at time of retrofire. Also

since the clock was not to be used for retrosequence, only

the minutes and seconds needed to be accurate for the astro-

naut's reference and the clock setting advanced 1 hour to

avoid possible clock timeout. This command was not set into

the clock, but the astronaut successfully reset the clock.

The telemetry and G-band beacon were again turned on for the

Range Tracker as the spacecraft passed over the CSQ.

The astronaut was advised to "take green for go, " which was

a coded means of telling him to take a dextro amphetamine

pill. The pill was used as an added precaution to be sure

that the astronaut was alert for the retrofire maneuver.

The flight surgeon was not concerned over the astronaut's

condition but he was not sure the astronaut was thoroughly

rested from his sleep. The ZZB site noted a rise in carbon

dioxide partial pressure (PCCu) and the astronaut was advised

to purge the suit circuit with fresh oxygen by going on

0 emergency rate flow. A quick estimate of the quantity ofd

1oxygen remaining indicated that Ip- hours of 0 emergency rate

flow was available. The ZZB capsule communicator confirmed

each item on the checklist with the astronaut and he verified

that all items on the list were completed with the exception

of arming the squib switch.

Two additional items were added to the checklist at this

time. The first was to be sure that the visor was closed,and the second was to cage the gyros.

For the first time, the astronaut reported to ZZB that the

main and standby 250 v-amp inverters had failed to operate.

At this point, MCC advised ZZB to turn the ASCS a-c switch

to off, because the inoperative power circuit would require

the entire reentry to be manually controlled.

The CSQ acquired communication with the spacecraft at 33:56:5^--

The capsule communicator and the astronaut reviewed the retro-

fire and reentry procedures and it was apparent that the

astronaut was prepared for his task. Time hacks were trans-mitted to the astronaut at retrofire time minus 60 seconds

and at retrofire time minus 30 seconds. Finally, a 10-second

terminal countdown to retrofire was transmitted. The squib

bus was armed at retrofire minus 5 seconds. The number one

retrorocket was ignited at 33:59:30 and ignition of retro-

rockets number two and three occurred at the proper 5-second



Page 8-11

intervals. The telemetry data immediately confirmed the

retrofire. The astronaut stated that his attitudes were

good and confirmed the ignition of the three retrorockets.

The retropackage was manually jettisoned at approximately

3 :00:43. Communications were good through loss of signal

with CSQ. The "beginning of the reentry blackout was re-

ported by the Range Tracker to be within 2 seconds of the

time predicted, which indicated that the landing point would

be close to nominal. Communications were regained by the

HAW capsule communicator via the relay aircraft at approxi-

mately 3 -:20:30. The weather in the recovery area was good.

The final landing point was only about 4 miles from nominal.

8.2.7 General Comments.- The network flight control teamsper-

formed extremely well. Communications between the ground

and the astronaut were concise and they conveyed thenec-

essary information. The flight control teams utilized the

proper contact and reporting procedures that were developed

for this mission. The network data as presented on the

summary messages appeared, in most cases, to have an accuracy

within 2 percent. The operations messages provided muchuse-

ful real-time data, and no difficulty existed in determining

the precise status of the spacecraft, the astronaut, and the

mission.

The ground communications were generally good; however, those

at the ZZB and CSQ sites were not as good as those from the

rest of the network. The air-ground communications as

monitored on the Goddard loop were somewhat better than in

previous missions and probably resulted from the use of

squelch control at the remote sites. The air-ground com-

munications as remoted through the HAW site were good;how-

ever, two transmitters at the HAW site failed.

The entire mission was an extremely smooth and well coordinated

effort. The preflight network simulation exercises were the

best planned and executed to date. The flight controllers

response was very good, and it is felt that this test was

the best executed mission or Project Mercury. The cooperation

between the flight astronaut and flight control personnel had

a significant influence on the success of the MA-9mission.

CQNriDEMTIAlr





1DENTIAL Page 8 - 13

absence of equipment malfunctions during the longer duration of the

MA-9 mission, considering that sites were allowed to go to a standby

status at various times, is indicative of a high degree of excellence

in maintenance and operational procedures by all sites.

8.2.1 Computing and trajectory displays.- The network countdown for

the MA-9mission began at the Goddard Computing Center at

midnight on May 14. The Goddard computer, equipment, interface,

CADFISS, and trajectory confidence tests were all satisfactory.

During the countdown, some dropout in the Goddard "B" computer

readouts was observed at the MCC. The high-speed output

subchannel on the data communication channel for this computer

was interchanged with the plotboard high-speed subchannel.

At the request of the Flight Dynamics Officer, the powered-

flight- phase was supported with the "A" and "C" computers,

and then the support was switched to the "A" and "B" computers

during orbital flight. The "B" computer gave no indication

of dropout during the rest of the mission. Lift-off occurredat 08:04:13 a.m.e.s.t.

The Atlantic Missile Eange (AMR) I.P. 709 and the General

Electric - Burroughs computers provided excellent data

throughout the launch. A "go"decision was indicated by all

three data sources. The cut-off conditions are shown in

table 8.2.1-1.

In the orbital phase, during the periods when the spacecraft

C- and S-band beacons were on, the tracking data received

from the network sites were excellent. During the mission,

weight changes in the spacecraft resulting from fuel andcoolant-water usage were manually put into the computers.

The retrofire time recommended by the Goddard computers was

33:59:30, and retrofire was manually initiated at this time.

After retrofire, the predicted landing point transmitted to

the MCC from the Goddard computer was 27° 22' North latitude

and 176° 29' West longitude. An attempt to refine this

prediction with six frames of data from the Range Tracker

ship, acquired during blackout, failed to yield a converged

solution. The computed time of the blackout was from

3 :08:16 to 3V.22:30. The actual time of initial blackout

was reported by the Range Tracker to be 3 -:08:17« The actuallanding point was reported by the recovery ship to be

27° 22.6' North latitude and 176° 35-3' West longitude.

Although several minor computer problems were encountered and

corrected throughout the flight, at no time during the mission

did the computers fail to drive the digital displays and

plotboards at the MCC. In addition, performance of the high-

speed lines between Goddard and the MCC was excellent.



3 -

For the first time, CADFISS tests were conducted during the

mission to determine the operational status of major equipment

subsystems at network sites. These tests were considered

necessary since mandatory equipment at many sites was shut

down for prolonged periods of time when the spacecraft was out

of range. All of these tests were successfully supported by

the third Goddard computer while the other two Goddard

computers continued the operational support of the mission.

This flight was the first to be supported by a triplex computing

system and the IBM 709 computers.

Two range ships, the Range Tracker and the Twin Falls Victory,

were used to provide tracking data to the Goddard computers.

The Range Tracker provided good tracking data during the 7th,

20th, and 21st orbital passes. During reentry the Range Tracker

was poorly positioned with respect to the blackout zone and

provided only six frames of data for this phase of reentry.

An analysis of these data indicated a landing point which was

about-3° or 180 nautical miles away from the correct landing

point. Twin Falls Victory data readout was good on three passes.

East Island data appear to be unusable. Ascension Island data,

used for the first time in a Mercury mission, appear to be

satisfactory.

8.2.2 Command system.- The command system for the MA-9mission

operated in a satisfactory mariner, and the command control plan

was followed very closely throughout the mission. Several

malfunctions were noted at various sites, but command capability

was never lost by any site during the time in which the space-

craft was passing over that site. The command carrier "on"

indication from BDA to the MCC was delayed approximately

32 seconds on the first pass; however, it had no net effect on

the mission since the onboard command receiver signal strength

remained above the receiver threshold setting. A summary of

the command handover exercises is shown in table 8.2.2-1 and

a summary of command transmissions is shown in table 8.2.2-2.

8.2.2.1 Ground system: A preliminary evaluation of the data shows

that all command sites had very good command coverage during

each pass. Command coverage became reliable at slant ranges

varying from 300 to 950 nautical miles. This large variancewas caused by the change in spacecraft antenna orientation

while the spacecraft was in attitude-free drifting flight, the

trajectory of"the spacecraft over the site, and the transmitter

output power of the respective sites.

MCC, GBI, and San Salvador (SAL) provided commandcon-

trol coverage during launch. This coverage was provided

by the MCC low pover unit from lift-off to T+90 seconds,

by the MCC high power unit from T+90 seconds to



Page 8-15

seconds, and by the SAL high power unit from

T+245 seconds to T+3&0 seconds, at which time command control

was switched to BDA. This switch to BDA was performed at

00:05:58.350, but the command carrier "on" indication was not

received at MCC until 00:06:30.3 , a delay of 32 seconds.

An evaluation of the onboard receiver signal strength data

revealed that the ground transmitted carrier was being received

by the onboard system at a signal strength of 27.5 microvolts

from lift-off to 00:07:03. The ground-recorded telemetry

readout correlates well with the onboard recorded data. This

delay is considered to have been a function of the ground

remoting equipment and was not detrimental to command control

during the launch phase. Investigations are continuing to

determine the cause of this delay.

The command control handover plan was updated one time during

the mission. This change compensated for the slight variance

in the orbital trajectory over the 22 orbital passes and was

sent to the Coastal Sentry Quebec (CSQ)ship for the reentry

phase. All other times were nominal.

A total of 19 functions was transmitted from the command

stations. All of these functions were received onboard the

spacecraft with the exception of one telemetry "on"function

from Muchea (MUG) and the clock change from the CSQ. The

telemetry "on"command from MUG was not received because it

was transmitted when the spacecraft was out of range of the

600-watt ground transmitter. The clock change from the CSQ

was not received because the command tone was also sent before

the spacecraft was within range of the ground transmitter.The frequency monitor aircraft which was stationed in the

Pacific reported through the CSQ that they did not attempt

any transmissions on the command frequency during the orbital

passes monitored.

The following ground system malfunctions were experienced:

1. The Bermuda high-power transmitter came on with a

3-6-kw output but did not come up to full power and failed

over to low power at approximately 00:06:16. This failover

did not interrupt command control during this pass. Beam

voltage was low and was corrected in time to support the nextpass with full 10 kw-power output.

"Vailover is defined as failure of the primary system,

accompanied by automatic switching to standby system.

•



8 - 1 6

2. The RKV had an intermittent problem in the beam

power supply of the backup power amplifier, and at 09:24:47 it

was reported as being inoperative for the remainder of the

mission. The prime transmitter was used to support the remainder

of the mission.

3. At BDA the C-band beacon "on" command was on before

the command carrier came up at 22:08:00, giving a code fault.

The code fault was released after the command system was brought

up manually.

4. Guaymas (GYM) had an autotransformer failure in the

standby transmitter at 25:16:47, and the system was reported

as operative at 27:13:47.

8.2.2.2 Spacecraft command system: The spacecraft had only one command

receiver onboard during the MA-9mission. The threshold level

was intended to be set between 2.5 and 3 microvolts; however,

the receiver was capable of and did receive command functions

at a level of approximately 5-microvolt. The saturation

value of the receiver was 27-5 microvolts. The system operated

normally with the exception of spurious command carrier reception

at 03:35:00 to 03:38:00, 11:24:00 to 11:27:00, and 27:10:40 to

27:13:20. During the time period of 27:10:40 to 27:13:20,

eight functions, having a duration of approximately 2 seconds

each, were recorded on the onboard recorder. These functions

have yet to be identified; however, they were of such a nature

that they did not affect the mission. An investigation is inprogress to determine the exact cause of these recorded events.

The preliminary results tend to show that these signals were

generated outside the spacecraft and were not caused by internal

EF beat harmonics.

The effect of antenna orientation angle during drifting flight

was noticeable, as it had been during the MA-8 mission.

However, the signal strengths received were much better than

had been anticipated. The onboard system was capable of and

did receive functions whenever they were transmitted within

range of the ground transmitters.

8.2.3 Eadar tracking performance.- During the countdown on May l4,

1963, the radar at Bermuda failed to pass the CADFISS slew

tests. Digital data were intermittently of poor quality in

both the azimuth and range channels. Efforts to locate the

trouble were ineffective, and the quality of the data

gradually decreased. At T-15 minutes, the range data error

exceeded the tolerable limits, and at T-13 minutes the mission

CONFIDENTIAL



Page 8-17

was postponed. Subsequent trouble shooting revealed a

faulty preamplifier in the azimuth digital data channel and

a faulty shift register in the range digital data channel.

The simultaneous failure of both components complicated the

failure analysis. The radar was repaired at 2:00 p.m. e.s.t.

on May l4, 1963.

On launch day there were no radar problems, and the C- and

S-band beacon checks prior to launch indicated no beacon

problems. The network C-band radars tracked approximately

10 percent of the total mission time, which is 80 percent of

the total time that the C-band beacon was turned on. The

network S-band radars tracked 1.7 percent of the total mission

time, which is 36 percent of the total time that the S-band

beacon was turned on. The amount of radar data furnished to

the Goddard computers was of sufficient quality and quantity

to update the trajectories, and it was determined that the

orbital parameters did not decay an appreciable amount.

Initial tracking reports indicated that the C-band beacon

was not as good as it had been on previous missions because

of the heavier than usual modulation on the beacon replies.

The heavy modulation experienced by the MCC and BDA radars

during launch seemed to lessen as the mission progressed.

In addition to the normal Mercury Network radar sites, the

following sites were used for the MA-9 mission: Ascension

Island, East Island, Puerto Rico, and the radar ships Twin

Falls Victory and Range Tracker. Radar tracking data for all

sites are tabulated in tables 8.2.3-1 and 8.2.3-2.

8.2.4 Ground telemetry system performance.- The telemetry coverage

for the MA-9 mission was excellent as shown in table 8.2.4-1.

There were no major ground system failures, although some

coverage was lost because of the manual switching procedure

used onboard the spacecraft. In general, any deviation from

nominal coverage can be attributed to spacecraft attitude or

to the transmitter being turned off. The telemetry relay

circuits from Antigua, California, Bermuda, and Ascension

were satisfactory in all respects. During all passes over

these stations when telemetry antennas were radiating, data

were remoted to the MCC. During the third orbital pass, the

telemetry was switched to the high-frequency link prior to the

spacecraft's passing over Hawaii and remained on until it

was over the California site, at which time telemetry was

switched back to the low-frequency link. At all other times,

the telemetry remained on low frequency. No telemetry system

anomalies were noted during this period.




In general, the performance of the acquisition-aid systems

at all stations was satisfactory and comparable to that of

previous missions. Low-angle elevation tracking, "below

approximately 15% was accomplished manually because of

multipath conditions at most stations. The only major

acquisition-aid problem experienced during the mission was

on the CSQ where failure of the elevation antenna drive

system occurred prior to the 6th orbital pass. However, theantenna was positioned manually from the 6th through the

8th passes, and the malfunction in the drive system was corrected

in time for acquisition in the 9~th pass.

8.2.5 Air-to-ground communications.- During the MA-9 mission, the

air-to-ground communications were of good quality. The UHF

system was used as the primary communications system except

for the scheduled HF checks. During periods of communication,

UHF coverage varied only slightly from predicted acquisition

and loss times because of the nominal orbital trajectory.

As expected, air-to-ground communications could not be

established during the communications blackout period. An

Instrumentation Support Instruction was transmitted to the

network outlining the use of the UHF squelch circuit as defined

in the network documentation. A premission checkout and the

mission results indicate that proper use of the squelch

circuit eliminates background noise from open UHF receivers

during periods of silence. This change also resulted in a

reduction of noise level on the Goddard circuit during

air-to-ground transmissions.

The results of the ground HF antenna test, in which a

vertically polarized antenna replaced the normally used

horizontally polarized two-element beam, are inconclusive at

this time. The Kano station reported that the signal strength

received was lower when the vertically polarized antenna was

used than when the horizontally polarized antenna was used.

Texas reported higher signal strengths when the vertically

polarized antenna was used during the second period of space-

craft contact. Since these reports are conflicting, analysis

of the test results is continuing.

Relay aircraft in the Atlantic Ocean area reported good UHFreception from the spacecraft and good relay transmissions to

MCC on the 2nd, 3rd, and 17th orbital passes. A relay attempt

on the l6th pass was unsuccessful because of a severe thunder-

storm in the vicinity of the relay aircraft. Communications

from the MCC to the spacecraft through the relay aircraft

were not attempted on the 2nd pass, and they were unsuccessful

on the 3rd pass because the spacecraft had passed out of range.

However, they were successful on the 17th pass. Ascension and



8 - 1 9

Antigua Islands in the Atlantic were also available for

relaying communications between the spacecraft and the MCC.

Relay through Ascension was successfully accomplished for a

period of approximately 6 minutes during the 3rd orbital

pass. The Antigua voice relay was not used during the

mission.

In the Pacific Ocean area, communications were successfully re-

layed from Hawaii through Kwajalein and ¥ake Islands on passes 3

and 19, respectively. A voice-operated relay from the MCC

through the Range Tracker was attempted on the 20th orbital pass.

However, this attempt was unsuccessful because the transmission

was made on the MCC/HAW remote air-ground position instead of

the Goddard Conference Loop. This error apparently placed a

1700-cps tone on the circuit to the Range Tracker and resulted

in keeping the auto-voice relay continuously closed; however,

several transmissions from the astronaut were received in the

MCC. Another attempt to use the relay on the 22nd pass wasineffective. As in the MA-8 mission, satisfactory communications

were established in the primary landing area between the space-

craft and Hawaii using relay aircraft.

Table 8.2.5-1 summarizes the air-to-ground communications

coverage.

Wfctfe



- 20

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Page 8-21

TABLE 8.2.2-1.- COMMAND HANDOVER SUMMARY

Orbital coverage: Mercury Control Center, 10 kw into quadhelix antenna;

Grand Bahama Island, 10 kw into sterling antenna; San Salvador, 10 kw

into sterling antenna; Bermuda, 600 w into quadhelix antenna (1st pass)

10 kw into quadhelix antenna (all other passes); Hawaii, California,

and Rose Knot Victor, 10 kw into quadhelix antenna; Muchea, Guaymas,

and Coastal Sentry Quebec, 600 w into quadhelix antenna

Station

Mercury Control Center


(San Salvador)

Bermuda

Muchea

California (backup to

Guaymas)

Guaymas


Bermuda

Muchea

Hawaii

California

Guaymas


Command carrier

On, g.e.t

Launch

00:04:05

(00:04:07)

00:05:58

(00:06:16)

00:45:00

(00:45:00)

01:23:00

(01:22:14)

01:20:00

(01:20:00)

01:33:00

(01:33:02)

01:38:00

(01:38:00)

02:15:00

(02:15:00)

02:45:00

(02:45:00)

02:56:00

(02:53:49)

03:04:00

(03:04:00)

03:06:00

(03:06:02)

Off, g.e.t.

00:04:05

(00:04:07)

00:06:00

(00:06:00)

00:12:00

(00:12:00)

00:59:00

(00:59:00)

01:30:30

(01:33:1 )

01:33:00

(01:30:00)

01:38:00

(01:38:00)

01:45:00

(01:45:00)

02: 32 : 00

(02:32:50)

02:56:00

(02:56:00)

03:04:00

(03:04:49)

03:06:00(03:06:00)

03:12:00

(03:11:50)

+10 [iv carrier

coverage above

line of sight,

percent

100

100

52

86

Mo radiation

85

100

97

95

100

93

30

91

xhe times given in parentheses are actual; times not given in parentheses

are planned.

Reduced signal strength and multi-path anticipated at the spacecraft due

to the low elevation angle and/or excessive slant range. Command Control Carrier

ON/OFF times are as noted.



- 22

TABtE 8.2.2-1.- COMMAND HANDOVER SUMMARY - Continued

Orbital coverage: Mercury Control Center, 10 kw into quadhelix antenna

Grand Bahama Island, 10 kw into sterling antenna; San Salvador, 10 kwinto sterling antenna; Bermuda, 600 w into quadhelix antenna (1st pass)




Station

Bermuda

tMuchea

Hawaii

California

Guaymas


San Salvador

Grand Turk Island

Hawaii

California

Guaymas


Hawaii

13

Command carrier

On, g.e.t.

03:12:00

(03:12:03)

03:54:00

(03:54:00)

04:15:00(04:15:00)

04:30:00

(04:27:27)

04:36:00

(04:36:00)

04:40:00

(04:40:02)

04:45:45

( O k - . h - y - . k 8 )

o4:47:oo(4:47:03)

05=56:00

(05:55=00)

06:03:00

(06:00:11)

06:09:00

(06:09:00)

06:14:15

(06:14:16)

07:30:00(07:30:00)

Off, g.e.t.

03:18:00

(03:18:00)

04:05:00

(Ok: 05:00)

04:30:00(04:30:00)

Ok : 36: 00(04:40:12)

04:40:00

(04:40:00)

04:45:45

(04:45:48)

04:47:00(04:47:02)

04:50:30(04:50:33)

06:03:00

(06:03:00)

06:09:00

(06:11:11)

06:14:15

(06:14:00)

06:19:30

(06:19:29)

07:35:00(07:35:00)

+10 p,v carrier

coverage atove

line of sight,

percent

100

90

69

84

100

100

100

97

100

100

100

100

77

The times given in parentheses are actual; times not given in parenthesesare planned.

Reduced signal strength and multi-path anticipated at the spacecraft due to

the low elevation angle and/or excessive slant range. Command Control CarrierOH/OFF times are as noted.

c o u n s e l



Page 8-23

TABLE 8.2.2-1.- COMMAND HANDOVER SUMMARY - Continued







Station

California13

Guaymas

Coastal Sentry Quebec

Hawaii


Hawaii


Hawaii13

Rose Knot Victor


Rose Knot Victor


Rose Knot Victor

Command carrier

On, g.e.t.

07:37:00

(07:34:28)

07:42:30

(07:42:30)

08:45:00(08:43:40)

09:02:00

(09:02:00)

10:18:00

(10:17:53)

10:35:00

(10:35:00)

11:53:00

(11:50:36)

12:10:00

(12:10:00)

12:23:00

(12:20:47)

13:25:00

(13:23:30)

13:55:00

(13:53:47)

15:00:00

(14:57:24)

15:30:00

(15:27:47)

Off, g.e.t.

07:42:30(07:44:34)

07:48:00(07:48:00)

08:54:30

(09:00:40)

09:10:30

(09:10:30)

10.28:00

(10:30:21)

10:45:00

(10:45:00)

12:02:00

(12:03:46)

12:17:00

(12:17:00)

12:33:00

(12:34:06)

13:36:00(13:37:42)

14:07:00

(14:07:45)

15:09:00

(15:10:53)

15:40:00

(15:41:03)

+10 p,v carrier

coverage above

line of sight,

percent

82

74

37

96

81

77

88

50

31

97

100

79

93

T?he times given in parentheses are actual; times not given in parentheses

are planned.




*



Page 8-24








Station


Rose Knot Victor

Rose Knot Victor

Rose Knot Victor

Grand Turk Island

Muchea


Bermuda

Muchea

Guaymas


Bermuda

Muchea

Command carrier

On, g.e.t.

(16:28:47)

17:04:00

(17:01:4?)

18:38:00

(18:34:47)

20:12:00

(20:08:47)

20:30:00

(20:30:02)

21:18:00

(21:18:00)

22:02:00

(22:02:01)

22:08:00

(22:08:00)

22:50:00

(22:50:00)

23:30:00

(23:30:00)

23:35:00

(23:35:01)

23:40:45

(23:40:45)

24:24:00(24:24:00)

Off, g.e.t.

(l6:42:0j)

17:14:00(17:14:42)

18:48:00

(18:48:44)

20:21:00

(20:21:58)

20:35:30

(20:35:34)

21:27:00

(21:27:00)

22:08:00

(22:08:00)

22:13:00

(22:13:00)

23:00:00

(23:00:00)

23:35:00

(23:35:00)

23:40:45

(23:40:46)

23:46:45

(23:46:00)

24:34:00

(24:34:00)

+10 | _ L V carrier

coverage above

line of sight,

percent

100

92

95

57

76

71

100

78

62

54

81

69

The times given in parentheses are actual; times not given in parentheses

are planned.






Page 8-25







_and Coastal Sentry Quebec, 600 w into quadhelix antenna

Station

California (backup to

Guaymas)

Guaymas


Bermuda

Muchea

Hawaii

California

Guaymas


Bermuda

Muchea

Hawaii

California

Command carrier

On, g.e.t.

25:00:00

(25:00:00)

25:00:00

(25:00:00)

25:10:00

(25:10:02)

25:14:00

(25:14:00)

25:57:00

(25:57:00)

26:23:00

(26:23:00)

26:32:00

(26:32:04)

26:39:00

(26:39=00)

26:43:00

(26:43:02)

26:48:00

(26:48:00)

27:30:00

(27:30:00)

27:56:00

(27:56:00)

28:05:00

(28:05:15)

Off, g.e.t.

25:07:00

(25:06:47)

25:10:00

(25:10:00)

25:14:00

(25:14:00)

25:20:00

(25:20:00)

26:07:00

(26:07:00)

26:32:00

(26:32:00)

26:39:00

(26:39:02)

26:43:00

(26:43:00)

26:48:00

(26:48:00)

26:53:00

(26:53:00)

27:40:00

(27:40:00)

28:05:00

(28:05:00)

28:12:00

(28:12:14)

+10 n,v carrier

coverage above

line of sight,

percent

1

95

55

86

88

100

56

64

5

100

100

97

73

T?he times given in parentheses are actual; times not given in parenthese

are planned.



OH/OFF times are as noted.



Page 8-26

TABI£ 8.2.2-1.- COMMAND HANDOVER SUMMARY - Concluded






and Coastal sentry Quebec* 600 w into quadhelix antenna

Station

Guaymas


Hawaii

California

Guaymas


. .bHawaii

California

Guaymas


Hawaii


Hawaii (backup for

Coastal Sentry Quebec -

reentry)

Command carrier

On, g.e.t.

28:12:00

(28:12:00)

28:16:30

(28:16:32)

29:30:00

(29:30:00)

29:38:00

(29:38:00)

29:1)4:30

(29:1(4:30)

(30:48:47)

31:05:00

(31:05:00)

31:13:00

(31:13:00)

31:15:00

(31:15:30)

32:22:00

(32:19:00)

32:38:00

(32:38:00)

35:55:00

(33:51:30)

34:11:00

Off, g.e.t.

28:16:30

(28:16:30)

28:23:00

(28:23:00)

29:38:00

(29:38:00)

29:44:30

(29:44:30)

29:50:00

(29:50:00)

(30:59: 3)

31 : 11 : 00(31:11:00)

31: 19 :~00

(31: 1 5 - A T )

31:22:00

(31:22:00)

32:30:'00

(32:32:19)

32:46:00

(32:46:00)

34:03:00

(34:05:30)

34:20:00

+10 u.v carrier

coverage above

line of sight,

percent

57

92

78

66

78

49

97

25

70

88

84

91

times given in parentheses are actual; times not given in parentheses

are planned.






Page 8-27

gH

O

I

O J

C M•

C X I

CO

C Q-p

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& 0-P rH

bD ofi -H

0) Be .rH

CO -P

H COC O fn

S) CL)•H 0w

&CO

C O*• CD

C O HM -Hfi BC OrH H

CO

-P 0S -H

C O -PH 302 CO

O

C Q

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aG o0 -H•H -P

C Q C O

CQ rH

13CQ

flC OrH 0

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CQi 1 " •

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ffl'gQ JJ p:•H ••CQ rH

"

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a >

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SO•H-POr_j

H

fi

C

0•H-P

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U A LT\ LtA UA OOO LfA UA UA

b- [>- t^ C O HA OJ IT l>- tC M O J C X J C M H C M C M C M

TdC O

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O O -3--3- C D O O QLP\OH rH O I<AVO O VO LT\-4" CM O O 0) t VO KA H fA

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L P \ VO -COO-^"LO\ -^HVOt -3F KA H rH

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L T \ KA NA f- tCOOJ OAOCMO \T\ O O LT\ UA O KA_4- -3"

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H H HC X I C X I C M

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H0rH-P Tjfn H

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L( A LtA LtA O

CXI t^- C^- t^-CM CXI CM

UA LT\ O LtACM O C^ KAO\ CMH VO

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CXI VO VD OALtA UA UA L!A

O J C M C M C MC X J CXJ CXI OJ

a a a a0 0 O O

T S Ti 'tj tjfi fi fi SC O C O C O C Oo o o o1 1 1 1

02 0CQ 0

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Page 8 - 2 8 • C O N F I D E N T I A L

TABLE 8.2.3-1.- COMPUTER READOUT OF RADAR TRACKING DATA

Station

1a

Bermuda,

Bermuda

Canary Islands

Canary IslandsMuchea

Woomera

White Sands

Eglin AFB

East Island


Bermuda

Hawaii

California

White Sands

Eglin AFB


San SalvadorEast Island

Bermuda

Ascension

Rose Knot Victor

Hawaii

Woomera


East Island

San Salvador

BermudaMuchea

Woomera

Eglin AFB

Pass no.

1

11

1

1

1112

2

2

3

333

1 *

k444

77141515

15

15151515

Total

7^7k6162

85

62

36702

67

775042

573

45

15564773

26

49584715

10

647466

59

Lines of data

Invalid

18

173311

38

9332

25

5228

71410

32

01

4T13

01833250

3

172

29

37

Garbled

0

01

01

000

00

2

40

0

0

0

00

0

0

0

00

0

0

0

000

0

Maximum elevation, deg

67-567-572.172.164.4

7 9 - 215.444.9

.718.2

96.025.036.580.936.6

37-2

84.447.72.825.0

59-0

71.770.2

8.86.9

2.0

14. 361.2

49.425.2

FPS-16

Verlort



Page 8 - 2 9

TABLE 8.2.3-1.- COMPUTER READOUT OF RADAR TRACKING DATA - Concluded

Station

Mercury Control CenterBermuda

HawaiiCaliforniaWhite Sands

Eglin AFBMercury Control CenterSan Salvador

East IslandBermuda

Rose Knot Victor

Rose Knot VictorHawaiiRose Knot Victor

Pas s no .

16

16

181818

18

19191919

20

21

21

22

Total

€ k6 3303 31 * 0

6831kl

839

2928

3720

Lines of data

Invalid

1624

2k

171

12

12

1

00

1

01814

Garbled

00

0

0

0

0

0

0

00

0

0

0

0

Maximum elevation, deg

65.0

76.521.1

27.6

70.2

28.0

27-3

56.9

3 3 - 02 - 3

59-848.59 - 1

24.1

C O N F I D E N T I A L..»_



8 - 30

T A B L E 8.2.3-2.- R A D A R T R A C K I N G P E R I O D S

StationDuration of signal

Acquisition, g.e.t. Loss, g.e.t.

C-"band

Mercury Control

Center

Grand Bahama

Island

San Salvador

East Island

Twin Falls Victory

Ascension

Bermuda

Woomera

00:00:0001:25:5404:43:36

22:05:59

28:17:45

00:00:52O i l - : 114:0028:18:47

00:02:0022:09:39

28:18:58

04:U6:0622:07:25

28:21:50

01:38:05

05:0

22:06:59

22:10:09

23: 0:33

Olf:59:2605:05:lU

00:03:25

01:37:38Ol4-:l| :39

22:06: 2

23:37: 52 :20:15

28:20:30

00:56:16

21:25:12

22:57:2

00:06:1001:itO:07

Oi4-:^7:32

22:09:13

28:23:21

00:05:0804:Vf:V7

28:23:30

00:06: 7

22:10:1+328:24:28

04:51:22:09:46

28:24:30

01:40:2004:48:07

22:08:3522:10:49

23:44:11

05:04:58

05:06:35

00:09:55

01:40:18

04:49:10

22:12:51

23:44:5124:20:27

28:23:41

01:02:4221:29:43

23:03=31



TABLE 8.2.3-2.- RADAR TRACKING PERIODS - Concluded

StationDuration of signal


C-band

Rose Knot Victor

Hawaii

California

White Sands

Eglin APE

10:33:2831:01:44

32:35:1034:09:51

04:23:40

05:57=1710:37:1727:58:14

31:06:20

32:40:02

01:27:05

04:29:06

28:ll:lU

01:29:5904:37:19

23:34:1*426:40:22

28:11:59

01:34:0104:39:21

23:37:28

28:15:16

10:36:2931:04:44

32:37:5434:10:33

04:28:3606:01:2710:43:2328:04:2231:10:16

32:45:37

01:31:0504:34:0928:13:17

01:33:22

04:41:21

23:37:0326:40:42

28:16:20

01:38:5404:4-5:57

23:40:16

28:21:25

S-band

Bermuda

Canary Islands

Muchea

00:04:11

00:15:47

22:19:41

00:49:45

22:53:27

00:09:57

00:21:35

22:23:29

00:58:07

23:00:44

C O N F I D E N T I A L... •* - .



Page 8-32

TABLE 8.2.4-1.- TELEMETRY COVERAGE

Duration of signal

Acquisition, g. e.t. Loss, g. e. t.Coverage., percent


0

01:33:2203:06:57

Ok: kO: 1806:15:06

22:04:15

23:36:5?25:09:20

26:42: 4528:16:10

29:50:37

00: 07: k201: 40: 42

03:14:15

Ok: 47: 1606:20:00

22:10:3723:44:21

25:16:2726:50:51

28:23:10

29:55:05

90

95959595

959095959595

Grand Bahama Island

00:01:14

01:34:1503:08:0004:41:30

06:15:20

22:03:50

23:37:00

25:10:30

26:44:30

28:17:30

29:51:50

00:06:44

01:40:1503:14:13

04:45:50

06:20:30

22:10:10

23:43:0525:16:45

26:50:3028:23:1529:55:50

98

98989898989898989898

* -ii "r ,



Page 8-33

TABLE 8.2.4-1.- TELEMETRY COVERAGE - Continued

Duration of signal

Acquisition, g. e.t. Loss, g. e.t.Coverage , percent

Grand Turk Island

00:02:30

01:37:0403:10:1204:43:00

06:17:0820:31:1022:04:09

23=38:45

25:13:0?26:46:08

28:18:5529:52=03

00: 07: 4201:41:10

03:15:0604:50:50

06:23:1020:37=3022:11:0423=43:4325:17:12

26:52:10

28:23:12

29: 58: 21

90100

100

95

95989595100

959895

Bermuda

00:03:12

01:36:4103:10:05

04: 44: 1520:35:38

22:06:31

23:39:36

25:12:5126:46:1228:20:26

00:10:2301:43:5603:15:0904:49:15

20:38:40

22:13:18

23:46:41

25:20:06

26:53=1228:23:14

95999999959999999995

O O M T D D M I A f c



TABLE 8. 2. 4-1. - TELEMETRY COVERAGE - Continued

Duration of signal

Acquisition, g. e.t. Loss, g. e.t.Coverage ^ percent

Canary Islands

00:14:1101:47:5217=37:1419:09:5420:44:0422: 17:34

23:50:3925:24:21

00:21:37

01: 54: 2717:43 = 3019:17:02

20:50:24

22:24:00

23:57:42

25=30=33

100

98

100

100

100

100

100

100

Kano

00:21:0901:54:4914:33:24

16:07:13

22:26:26

23=57:25

25:31:15

00:28:2902:01:30

14:40:23

16:13:13

22:28:21

24: 04: 4225:37:20

98979898949898

OONnDDPiTIAL1



Page 8 - 35

TABLE 8. 2. 4-1. - TELEMETRY COVERAGE - Continued

Duration of signal

Acquisition, g. e.t. Loss, g. e.t.Coverage, percent

Zanzibar

00:29:5902:04:02

10:00:0211:52:5622:35:2?

23=53:5525:40:42

33: 32: 42

00:39:1702:11:1210: Ok: 0411:38:0522: 38: 4224:02:02

25:46:52

33=39:32

9690888995979697

Muchea

00:49:1902:22:52

03:58:1221:19:40

22:52:1224:25:30

25:58:5527:33:52

00:58:1102:31:40

04:04:2721:27:1123:00:50

24:35:1226:07=3727:40:06

9697979794979898

C O M r i D D W T I A L



Page 8-36

TABLE 8.2.4-1.- TELEMETRY COVERAGE - Continued

Duration of signal

Acquisition, g. e.t. Loss, g. e.t.

Coverage } percent

Canton Island

01:09:2502:43:10

12:16:15

13:50:14

24:11-5:2526:19:21

01:l6:4l02:49:45

12:24:06

13:57:0324:53:01

26:25=35

989995

989799


08:52:00

10:21:30

11:55:0?13:28:34

15:02:02

16:37:4532:23:26

33:56:52

08:54:5710:28:22

12:01:4613:35:4215:08:03

16: 40: 0332:30:19

34:03:30

4090

95100

95059890



. C O i n T D D O T f A L Page 8-37

TABLE 8.2.4-1.-- TELEMETRY COVERAGE - Continued

Duration of signal

Acquisition, g. e.t. Loss, g. e.t.

Coverage .percent

Hawaii

02:49:1704:22:01

05:57:20

07:30:5109:04:0210: 37: 0312:11:1826:24:50

27:57:5729:35:0231:07:00

32:39:3

02:55:12

04:29:0206: 01: 45

07:35:0709:10:25

10:44:48

12:17:1726:31:22

28:04:3229:37:20

31:10:4132:45:50

809392

106075759075958580

Rose Knot Victor

12:25:26

13:58:22

15:32:2417:06:30

18:39:5720:13:34

12:33:20

14: 07: 14

15:40: 4017:14:2218:48:3120:21:14

989898939898



Page 8-38

TABLE 8.2.4-1.- TELEMETRY COVERAGE -Continued

Duration of signal

Acquisition, g. e.t. Loss, g.e.t.

Coverage, percent

Guaymas

01:25:57

03:00:2304:34:22

06:07:33

07:40:59

23:30:23

25:02:4226:36:2828:10:03

29:43:18

31: 16: hO

01:33:27

03:06:4704:40:2706:14:1707:47:20

23:35:3725:09:31

26:42: 4728:16:1729: 50: 0331:22:40

80

89969694949795969894

California

01:26:4?02:58:3704:31:0706:04:0207:38:17

25:02:3226:34:27

28:06:27

29:39:0731:13=47

01:31:07

03:05:07

04:38:32

06:11:0307:44:17

25:07:32

26:41:02

28:14:2529:47:17

31:19=47

95989795989899989798

C O I U r i D D N T I A t r



Page 8-39

TABLE 8.2. -1.- TELEMETRY COVERAGE - Continued

Duration of signal


Coverage, percent

Texas

01:29:31

03:03:20

0 :36:52

06:10:0807:50:00

22:02:05

23:32=2225:05: 0

26:39:25

28:12: 529: 6:56

31=12:35

01:36:25

03:09:550 : 3:55

06:17:30

07::07

22:03: 7

23=39:1^25:12:3

26: 5:5528:19: 229:53:00

31:23:20

70100

100

100

100

100

100

9 8100

90100

50

Ascension

06:3 :39

12: 52: Ml-

14: 2k: 316:02:13

27:01: 0

06:39:3812: 58: 571 :33:21

16:02:17

27:09:39

9999999999

C O M r i D D P i T I A L



Page 40 • C O N F I D E N T I A L

TABLE 8.2.4-1.- TELEMETRY COVERAGE - Concluded

Duration of signal

Acquisition, g. e. t. Loss, g. e.t.Coverage, percent

Antigua Island

05:12:5304:45:22

06:19:5218:59:40

20:32:10

22:07:00

26:48:5028:21:20

29:55:50

03:13:3004:52:30

06:25:2419:05:20

20:39:30

22:11: 40

26:54:3528:23:1530:00:29

989898989898989898

Pretoria

02:05:50

06:44:25

27:12:37

31: 53: 2033:26:46

03:10:3706:52:20

27:18:2632:02:07

33:33:00

9898989898

fflFIDEMTIAL-



Page 8 - 4l

TABLE 8.2.5-1.- AIR-GROIMD COMMUNICATIONS COVERAGE

OrbitalPass

Duration of signal


Average UHF signal

strength, microvolts


1

2

34515161718

1920

00:00:00

01:33:5303:OT:2904:40:3006:15:28

22:02:30

23:37:0025:11:24

26:45:50

28:16:38

29:52:00

00:10:3001: 42 : 04

03:14:01

04:49:20

06:25:20

22:12:28

23:43:1525:17:52

26:47:3828:21:28

29:56:00

1000

40

345352303T501008048

Grand Bahama Island

12

34

51516

IT181920

00:00:27

01:34:22

03:08:0704:41:1706: 16: 17

22:03:3T23:36:22

25:11:3226 : 45 : 42

28:l6:5T

29:50:07

00:07:1701:40:12

03:13:0704:48:12

06:19:1722:09:04

23 : 43 : 07

25:16:3226:47-:4l

28:21:25

29:55:06

(a)

Grand Turk Island

1

2

3

5

1516IT18

1920

00:03:18

01:37:1303:10:10

04:43:18

06:17:13

22:04:1823:39:18

25:13:0326:46:4828:18:48

29:52:48

00:06:48

01:41:0303:13:48

04:49:13

06:20:03

22:10:4323:43:38

25:16:43

26:48:4828:20:48

29:55:48

(a)

Recordings not available



Page 8 - 4 2

T A B L E 8.2.5-1.- A I R - G R O U N D C O M M U N I C A T I O N S C O V E R A G E - Continued

Orbital

Pass

Duration of signal


Average UEF signal


Bermuda

12

34

1516171819

00:03:29

01:36:4803:10:27

04:44:39

22:06:52

23:39:5925:14:00

26:46:23

28:21:28

00:09:3301 : 42 : 0403:14:0104:48:1322:12:28

23:43:15

25:17:52

26:47:38

28:21:30

41

702812

17-5804l

29(*)

Canary Islands

12

1516IT

00:15:10

01:48:3522:18:24

23:51:3725:25:54

00:21:20

01:50:23

22:23:31

23:56:1025:30:09

5.4l5 - 3

(c)

3 - 5

Kano

1

2

1516

17

00:22:28

01:55:0922:27:09

23:57:4825:31:49

00:27:48

01:55:1422:27:54

23:59:3125:35:17

3117-5624

11

Muchea

12

31415161718

00:51:0702:24:17

03:58:3721:22:27

22:53:12

24:27:36

26:00:39

27:33:23

00:56:5702:30:47

04:04:1321:26:29

22:57:20

24:33:1326:07:14

27:40:00

6038

(c)4o

25605022

bWo contact

'No record



Page 8 - 4 3

' TA B LE 8.2.5-1.- A I R - G R O U N D C O M M U N I C A T I O N S C O V E R A G E - Continued

Orbital

Pass

Duration of signal


Average UEF signal

strength^ microvolts

Canton Island

12

910

16

17

01:09:25

02:43:11

13:44:3715:18:4024:45:22

26:19:14

01:16:42

02:49:49

13:52:40

15:25:3224:53:0326:25:40

22

20

(b)(b)22

5

Guaymas

12

34

515161718

1920

01:27:1603:01:00

04:35:22

06:08:42

07:4l:5423:30:50

25:03:0726:36:41

28:12:49

29:43:46

31:16:43

01:29:1003:04:17

04:37:17

06:10:34

07:45:5823:33:16

25:09:1326:36:43

28:13:18

29:49:18

31:22:39

254o

2530

151250(c)10

1515

Zanzibar

12

78

1516

1722

00:30:5502:05:23

09:59:511:32:29

22:36:14

24:07:24

25:41:37

33:32:55

00:37: 5

02:19:53

10:03:3711:36:02

22:37:49

24:14:04

25:42:43

33:38:47

32.2

1 3 - 5185521

26.9

19(c)

Texas

2

31 *

51618

19

03:03:47

04:36:5506:16:22

07:45:47

25:06:12

28:13:02

29:46:12

03:10:02

04:43:57

06:17:3707:48:17

25:12:44

28:19:52

29:53:12

12

20

15204020

25

Wo contact

"No record



Page 8 - 44 E K T I A L

TABLE 8.2.5-1.- AIR-GROUND COMMUNICATIONS COVERAGE - Concluded

Orbital

Pass

Duration of signal

Acquisition, g.e.t. | Loss, g.e.t.

Average UHF signal



56

18

92021

22

07:17:20

08:51:3710:21:30

11:55:01

13:28:34

30:50:4632:22:31

33:56:5

(a )08:5 :5710:28:22

12:01:46

13:35: 2

30:51:5632:31:06

34:03:18

(a)

Rose Knot Victor

8" 12:25:47 12:33:32 18

Hawaii

2

34

56

7

81718

1920

21

02:51:4704:22:12

05:58:37

07:31:2509:0 :17

10:37:33

12:10:1526:26: -5

28:00:11

29:33: 7

31:07:1732:40:22

02:52:17

04:25:12

06:05:15

07:3 :50

09:09:52

10:43:40

12:14:2026:26:54

28:01:47

29:36:42

31:11:47

32:45:57

34110

722

29

51

374880

191918

California

12

3

4516

1718

1920

01:28:22

03:00:02

0 :33:16

06: 06: 0007: 0:5525 : 03 : 2226:36:32

28:07:5129:41:17

31:16:22

01:29:22

03:05:02

04:37:02

06:11:4707:44:0?25:07:27

26:38:12

28:13:27

29:44:32

31:19: 5

20

1624

2514

1514

393^30

Recordings not available

One contact only



EH~* Page 9-1

9.0 RECOVERY

9-1 Recovery Plans

The areas where recovery ships and aircraft were positioned in the

Atlantic and Pacific Oceans are shown in figure 9-l-l(a) and figure

9-l-l(b), respectively. Recovery capability was provided in areas A

through F in the event that it became necessary to abort the mission

during powered flight. Areas 2-1 through 21-1 were areas in which the

spacecraft could have landed if the flight were terminated earlier than

planned. These areas were spaced so that the spacecraft would pass

over one of them approximately every 90 minutes, or about once per orbital

pass. Area 22-1, which was the primary planned landing area for a nom-

inal flight of 3 - hours and 20 minutes, was located approximately 70

nautical miles southeast of Midway Island.

Recovery forces were deployed within these planned landing areas so"

that recovery and assistance could be provided within 3 to 9 hours after

spacecraft landing. This "access time" varied for the different areas

and was based on the probability of a spacecraft landing within a given

area and the planned deployment of recovery forces in that area. Selec-

tion of landing areas at spacecraft ground-track intersections permitted

a unit to move from one area to another and thereby provide a recovery

capability in several landing areas. A total of 23 ships and kk aircraft

were employed in the MA-9 recovery operation, of which 12 ships and 26 air-

craft were in the Atlantic landing areas and 11 ships and 18 aircraft were

in the Pacific. Table 9-1-1 indicates the number of ships and aircraft

on station at the various landing areas, their movements from one area toanother, and the access time for each area. Additional search aircraft

were available as back-ups to the aircraft on station. Also, helicopters,

amphibious surface vehicles, and small boats were positioned for recovery

support near the launch complex.

Contingency recovery aircraft and personnel were on alert status at

staging bases around the world to provide support in the event a landing

should occur at any place along the orbital ground track. These aircraft

were equipped to locate the spacecraft and to provide emergency on-the-

scene assistance if required. A typical support unit at a staging base

consisted of 2 or 3 long-range aircraft and pararescue personnel.

The locations of these staging bases are as follows:

Patrick Air Force Base, Florida Singapore, Malaya

Kindley Air Force Base, Bermuda Clark Air Force Base, Philippines

Lajes Air Force Base, Azores Waha, Okinawa

Nouasseur, Morocco Tachikawa, Japan

Wheelus Air Force Base, Libia Andersen Air Force Base, Guam

Kano, Nigeria Perth, Australia

Aden Protectorate Townsville, Australia

C O M F 1 D E M T M L



Page 9-2

Nairobi, Kenya Midway Island

Salisbury, Rhodesia Kwajalein Island

Mauritius Island Wandi, Fiji Islands

Ascension Island Hickam Air Force Base, Hawaii

Trinidad Island Johnston Island

Lima, Peru Papeete, Tahiti

Galapagos Islands San Diego, California

9-2 Recovery Operations

All recovery forces were on station at launch time and moved to

planned positions as the mission progressed. Weather conditions were

favorable for spacecraft location and retrieval in all planned andcon-

tingency areas throughout the mission.

A communications network linked the deployed recovery forces to the

recovery room of the Mercury Control Center at Cape Canaveral. Recovery

communications were good throughout the entire operation and the recovery

forces were informed of mission status during all phases of the flight.

At an appropriate time, recovery units were informed that the flight

would proceed to normal completion, and the expected retrorocket ignition

time of 33:59:30 for landing area 22-1 was transmitted. At about 34:04:00,

approximately l6 minutes prior to landing, recovery forces in area 22-1

were informed that the retrorockets had ignited normally and that the

landing position was predicted to be 2T°23rNorth latitude and 176°30' West

longitude. This information was transmitted as CALREP 1 (calculated land-

ing position report) to the recovery forces from the Recovery Coordinator

in the Mercury Control Center.. Recovery units in the area made contact

with the descending spacecraft before any additional predicted landing po-

sitions, based on reentry tracking, were made available from network support.

At about 34:12:00 the U.S.S. Kearsarge, the aircraft carrier positioned in

the center of area 22-1, reported radar contact with the spacecraft at a

slant range of l80 nautical miles and held contact until shortly before

spacecraft landing. A "sonic boom," similar to that heard during the

MA-8 reentry, was detected by recovery ship personnel. Personnel aboard

the U. S. S. Kearsarge reported first visual contact with the spacecraft as

it descended on the main parachute at an altitude of about 8,000 feet.

(See fig. 9.2-1.) The spacecraft landed at approximately 34:20:00 at27°22.6' North latitude, 176°35.3' West longitude, which corresponded to

a position approximately 4.4 nautical miles uprange of the recovery ship.

Weather conditions in the recovery area are listed in section 10.3.

Helicopters had been launched from the U.S.S. Kearsarge 15 minutes

prior to spacecraft landing and were in an excellent position to deploy

swimmers immediately. These swimmers quickly installed the auxiliary

flotation collar around the spacecraft. Helicopter pilots and swimmers

1M.J\Li



Page 9-3

in the landing area reported that the spacecraft, Immediately after land-ing, was floating on its side and gradually righted itself soon after mainparachute release, which was estimated to have occurred about JO secondsafter landing. Four minutes after landing, the astronaut reported thathe would remain in the spacecraft and await retrieval "by the recovery ship.

The antenna canister landed within 600 feet of the spacecraft but sankbefore the back-up swimmers could attempt retrieval. The ejected reserveparachute was retrieved by the swimmers. At about 5k:k6:OQ as the U. S. S.Kearsarge approached within 600 feet of the spacecraft, a motor whaleboat(shown in fig. 9.2-2) attached a lifting line to the recovery loop of thespacecraft. The spacecraft was then brought alongside the ship, liftedclear of the water, and placed on the no. 3 elevator of the recoverycarrier at 3 t56:00. The explosive -actuated hatch was released, as shownin figure 9-2-3, by using the external release lanyard. At about 35sOlsOO>doctors began examining the astronaut and taking blood-pressure measure-ments. The astronaut egressed from the spacecraft kQ minutes after landingat 35:08:00. (See fig. 9.2- .) He remained onboard the Kearsarge for a

period of examination, rest, and debriefing.

There was no apparent damage to the spacecraft at the time of landing.The swimmers who attached the auxiliary flotation collar to the spacecraftreported that none of the heat -shield straps were broken and damage tothe landing bag consisted of several small vertical tears. However, whilethe spacecraft was being lifted aboard the carrier, the UHF descent andrecovery antenna was broken loose at the hinge point. In addition, whenthe explosive -actuated hatch was released, the spacecraft window wasbroken.

Certain spacecraft onboard equipment was removed immediately after

recovery and flown to Cape Canaveral. Two days after recovery, the space-craft was transferred from the recovery ship to a truck at Pearl Harbor NavalBase. Prom there, it was taken to Hickam Air Force Base, loaded aboard aC-130 aircraft, and delivered to Cape Canaveral the next day.

9 - 3 Recovery Aids

Prior to spacecraft landing, telemetry aircraft established contactbefore and after communications blackout, and radar tracking aircraftmaintained contact during most of the blackout period.

All spacecraft visual and electronic recovery aids were reported tohave been operating normally. Helicopters reported that the dye markerpresented a brilliant green visual target which could be observed from adistance of 500 yards. The flashing light was reported as operatingnormally by on-scene observers.



p a g e 9 - ooinrroraiTLiL

The HF/DF- stations in Midway Island and San Francisco, California,

received, spacecraft • HF transmissions after landing and determined the

landing position within approximately JO nautical miles of the actual

retrieval point, as shown in figure 9- 3-1-

The telemetry and search aircraft reported contact with and verifiedthe operation of both spacecraft SARAH beacons. Acquisition ranges re-

ported by these aircraft were as great as 270 nautical miles.

Stations reporting fixes from the SOFAR-bomb detonation determined

the landing position within 10 nautical miles. A quick fix was provided

approximately 20 minutes after landing. Post-landing details of area 22-1

are illustrated in figure 9-3-1-



Page 9-5

TABLE 9.1-1.- RECOVERY SHIP AND AIRCRAFT DEPLOYMENT IN PLANNED LANDING AREAS

AreasAccess

time, hoursShips

Aircraft

Cb)

Atlantic

A

BCD

E

F

2-1

16-1 and 17-1

3-1 and 18-1

15-1

1*41

13-1

69636633

333

5

3 destroyers

4 destroyers and 1 carrier

1 destroyer

1 fleet oiler

1 fleet oiler

1 destroyer

1 destroyer, 1 carrier from area B

Area 2-1 ships plus additional

destroyer from B

2 destroyers from A

2 destroyers (l from B, 1 from C)1 fleet oiler, 1 destroyer from D

and E

1 destroyer from F

2

311113

9k

2

21

Total 12 26

Pacific

9-110-1

11-1

3-2, 8-1, and

18-2

-1, 7-1, and

22-1

5-1 and 6-1

20-1 and 21-1

553

3

33

1 destroyer

1 destroyer

2 destroyers

2 destroyers

2 destroyers, 1 carrier

2 destroyers

2

2

2

2

2

122

Total 11 18

Carriers had recovery helicopters embarked.

Numbers indicate maximum number of aircraft supporting a specific

area at the calculated time of spacecraft landing in that area. In somecases aircraft supported more than one area.

C O M T D D P i T I A L



P a g e 9 - 6

A i



P a g e 9- 7





Page 9

..

IJy

;

j.

•m•**'

t K . ^



C O N F I D E N T I Page 9-1




Page 9-11



Page 9-12

S a p '



Page 10-1

10.0 APPENDIX A

10.1 Spacecraft History

Spacecraft 20 arrived at Cape Canaveral, Florida, on October 9>1962. The preparation period for the spacecraft and the individualonboard systems and equipment was slightly different from that ofprevious spacecraft in that an integrated systems test was made priorto the normal individual system tests. This sequence of testing waspossible because of the extensive individual spacecraft systems testswhich had been performed by the spacecraft contractor. The number

of work days in the hangar totaled 170 days, of which 59^ days were

spent on formal tests. There were 6ll Mission Preparation Sheets (MPS),

which authorize specifically required work, and 62^ DiscrepancyReports (DR), which describe items requiring rework.

The spacecraft was transported to the launch site and mated with

the launch vehicle on April 22, 19&3-

The major prelaunch tests, modifications, and events in thehistory of spacecraft 20 at Cape Canaveral are shown in chronological

order in the following table:

A ¥





Page 1 0- 3

Number

17

18

19

20

21

22

23

2U

25

26

27

28

29

30

3 132

33

Item

Temperature survey wiring removed

Recovery light manual disable svitch

installed

HO low-pressure indicators installed

Thruster "B" nut temperature wiring

installed for T/M and recorder

Geiger counter calibration completed

CO adsorber assembly insulation

installed

Coolant control valve flow rate tests

were completed

BPMS microdot cable was replaced with

a seven- strand cable

A later model pitch attitude gyro

having a -3 -° pitch caging

capability was installed

Final inspection of TV circuitry

Television system test completed

Communications system radiation test

completed

Trial mating of spacecraft with

adapter

Main clamp-ring fitted

Air deflector installed on cabin fan

Urine bags and couch installed

Primary simulated flight in hangar

Completion date

Feb. 2, 1963

Feb. 3, 1963

Feb. 5, 1963

Feb. 10, 1963

Feb. lU, 1963

Feb. 16, 1963

Feb. 18, 1963

Feb. 20, 1963

Feb. 20, 1963

Feb. 28, 1963

Mar. 1, 1963

Mar. k, 1963

Mar. 5, 1963

Mar. 7, 1963

Mar. 8, 1963

Mar. 16, 1963

Mar. 23, 1963



Page 10 - I V F I D E P r T L I L

Number Item Completion date

35

36

37

38

39

i n

50

Adapter transported to the launch site

Automatic stabilization and controlsystem (ASCS) test completed

Low-level fuel warning indicatorsystem installed

ECS system reworked to flight config-uration. The automatic and reservefuel tanks paralleled.

Special test to evaluate inverter out-

put voltages completed

Prepad RCS test completed

Special screen adde d to the negativepressure relief valve

N -pressurized whip antenna installed

Flight batteries installed

Final heat-shield drop tests

Final prepad cabin-leak check(i+50 cc/min)

Alinement, weight, and balancecompleted

Spacecraft transported to the launchsite for mating

Simulated flight no. 1

Electrical mate

Simulated flight no. 2 (Joint FACT)

Urine transfer pump installed

Mar. 26, 1963

Mar. 26, 1963

Mar. 27, 1963

Mar. 28, 1963

Apr. h, 1963

Apr. 6, 1963

Apr. 10, 1963

Apr. 1 * 4 - , 1963

Apr. 15, 1963

Apr. 16, 1963

Apr. 17, 1963

Apr. 18, 1963

Apr. 22, 1963

Apr. 23, 1963

Apr. 2U, 1963

Apr. 25, 1963

Apr. 30, 1963



K F I D E K T I A L Page 10-5

Number

51

52

53

5U

55

56

57

58

Item

Simulated flight no. 2., joint FACT(repeated)

RCS checks completed

Launch simulation

Condensate trap installed in the suit

circuit

Simulated flight no. J

One of the squibs in the retropackage

explosive bolt disabled

Launch postponed because of radar

encoder problem at Bermuda site

Final launch countdown and lift-off

Completion date

May 1, 1963

May 1, 1963

May 8, 1963

May 9, 1963

May 10, 1963

May 13, 1963

May 'Ik, 1963

May 15, 1963

M F I D E P O I A L



1 0 - 6 • C O N F I D E N T I A L

10.2 Launch Procedure

The space-vehicle launch operations were planned about a 560-minute

split countdown, with a scheduled 19-hour hold at T-390 minutes for

fueling of the spacecraft reaction control system and servicing of pyro-technic systems. To provide additional assurance that the projected

launch time of 8:00 a.m. e.s.t. could be met, a scheduled 90-minute

hold was established at T-l4o minutes.

The second half of the split countdown was started at midnight on

May 13, 1963- Because of a fuel-pump failure in the gantry diesel

engine, a hold lasting 2 hours and 9 minutes was begun at T-60 minutes

to repair this malfunction. The countdown was resumed, but at T-13

minutes, another hold was called to evaluate a malfunction in the C-band

encoder at Bermuda-. The nature of this problem prompted a decision at

9:56 a.m. e.s.t. to postpone launch operations until this problem could

be corrected. The flight was rescheduled and the second half of thesplit countdown was started at midnight on May lk, 1963. The countdown

proceeded smoothly; and after a ^--minute hold at T-ll minutes, launch

occurred at 8:0 -:13 a.m. e.s.t. on May 15, 19 3-

The following is a sequence of major events which occurred during

the final countdown:

Time, min

T-390 Start of second half of countdown

T-l4o Astronaut insertion

T-107 Spacecraft hatch closure began

T-92 Spacecraft hatch closure secured; shingle

installation began

T-82 Spacecraft shingle installation complete

T-62 Service tower removal

T-35 Lox pumping started

T-24 Lox filling complete

T-ll A ^--minute hold was called to evaluate an external

KF interference problem with the guidance central

rate station.

T-0 Lift-off



Page 10-7

10.3 Weather Conditions

Weather conditions in the area of the launch site vere completely

satisfactory for operations several days prior to and on the days of

the MA.-9 flight. However, there was a trough in the Atlantic Oceannear the Caribbean which was moving north and could have caused concern

in some of the mid-Atlantic recovery areas had launch been postponed

longer than a few days. In addition, a weather front was moving down

the Atlantic seaboard that could have resulted in unsatisfactory launch

conditions, but these conditions did not materialize.

In the western Pacific Ocean, a tropical disturbance near the

southern Philippine Islands did not intensify sufficiently to cause

problems in the planned recovery area. Possible scattered light

showers and cloud cover were the only items of concern for the

Pacific recovery areas, and conditions were deemed satisfactory on

the day of launch for normal recovery operations.

During the 2 days of flight, exceptionally good weather conditions

were prevalent around the world. High-pressure regions which were

prevalent throughout the latitudes of the orbital ground track resulted

in only slight cloud cover around the entire ground track and in excel-

lent visibility for the astronaut from orbital altitudes.

Weather observations in the launch area at 8:07 a.m. e.s.t., just

after lift-off, were as follows:

Wind direction, deg 270

Wind velocity, knots 5Temperature, °F 7 .8

Relative humidity, percent J2

Dew point, °F 65

Visibility, miles 8

Cloud coverage (Cirrus with haze aloft) 5/10

Pressure, in. Hg 30.095

A plot of the launch-area wind direction and speed is shown in

figure 10.3-1.

K r i D E P F T I A L



Page 10 - C O N F I D E N T I A L

The weather and sea conditions reported in the primary Pacific

landing area by the Kearsarge at landing were as follows:

Wind direction,, deg 060

Wind velocity, knots 19

Wave direction, deg 060Wave height (at 6 sec intervals), ft

Swell direction, deg 080

Swell height (at 9sec

intervals), ft

Cloud cover (l,500 ft scattered; 10,000

ft scattered; high, broken) 8/10

Relative humidity, percent 69

Temperature, °F 75

Sea temperature, °F 75

Wind shifted from 090° at 10 minutes prior to landing.



Page 10-9

10.h Flight Safety Review

Flight safety and mission review meetings were conducted to determine

the flight worthiness of the MA-9 spacecraft and launch vehicle and to

ascertain the readiness of all supporting elements for the MA-9 mission.

10.h.I Spacecraft.- As a result of the systems changes necessitated by

the extended mission, flight safety review meetings were held

after the completion of the RCS and ECS systems tests and after

the primary simulated flight in the hangar. The meeting on the

RCS system was held on November 29, 19 2, and, for that stage of

checkout and preparation, the system was found to be capable of

performing,'its intended mission. The meeting on the ECS system

was held on February 5, 19&3, and no problems other than those

already in the process of being resolved were noted. Again, for

that stage of checkout and preparation, the system was found to

be capable of performing it's mission. A review of the spacecraft

systems following the primary simulated flight was held onApril 9, 1963?

arid all problems encountered were discussed and

direction for corrective action was given. At the spacecraft

flight safety review held on May 9, 196j, all systems were

approved as ready for flight, pending the successful completion

of the final simulated flight test, which was satisfactorily

completed on May 10, 1963.

10. U. 2 Launch vehicle.- Several meetings were held to determine the stat

of the Atlas IJOD launch vehicle. The first meeting was conducted

on April 22, 1963, to review the nature and the solutions of the

problems in the flight control system that had caused the late

delivery of the launch vehicle to the Atlantic Missile Range.Satisfactory resolution had been obtained on all problems. The

regular launch-vehicle review meeting was held at 9'00 a.m. e.s.t.

on May 11, 1963- The status of the launch-vehicle systems was

reviewed and all systems were approved for flight, pending the

resolution of a ground-equipment fuse anomaly. Subsequent testing

revealed the anomaly to be an out-of-tolerance fuse, and this com-

ponent was replaced.

10. -3 Mission.- The MA-9 mission review meeting was held on May 11, 1963-

The launch vehicle was listed in a no-go status because of the fuse

problem mentioned in the preceding paragraph. The spacecraft was

also determined to be in a no-go status pending a leak check on a

pressure transducer in the reaction control system. In addition,

trouble with Bermuda tracking was reported. All other elements of

the flight were found to be ready.

- C O N F I D E N T I A L



P a g e 10 - 10 G O F f r i D D H T I A L

The X-l Day Flight Safety Reviev Board met on May 1J, 1963. This

board was advised that the Launch-Vehicle Status Reviev Board had

met earlier at 8:15 and had determined the Atlas 1JOD was ready

for flight. In addition,, problems with the spacecraft and

Bermuda tracking were reported as having been solved, and the

Flight Safety Review Board approved both the MA-9 spacecraft andlaunch vehicle for flight.



1 0 -

10.5 Photographic Coverage

Photographic coverage, including quantity of instrumentation com-

mitted and data during the launch phase, which was obtained "by the

Atlantic Missile Range (AMR) is shown in table 10.5-1 and discussed

in the following paragraphs. Additional coverage obtained by the re-covery forces in the landing area is also described. Launch-phase photo-

graphic coverage was generally poor in quality because of sunlight and

overcast haze conditions. However, photographic data were obtained

through the time of launch-vehicle staging and were available for a

detailed photographic evaluation had it been necessary. The photo-

graphic coverage discussed in the following sections is based on film

available for evaluation during the postlaunch reporting period.

10.5-1 Metric film.- Metric film from 16 cameras were processed,

and the results were tabulated by the AMR. These data were

not required for evaluation by the Manned Spacecraft Center,

since the powered-flight phase was normal.

10.5-2 Engineering; sequential film.- Engineering sequential coverage

of the launch phase is shown in figure 10.5-2-1. This figure

indicates the time interval for which the spacecraft, launch

vehicle, and/or exhaust flame were visible to the tracking

camera. Optimum camera coverage was obtained from lift-off

through the region of maximum dynamic pressure, and adequate

data would have been available had a malfunction occurred

during this time. Although photographic coverage was obtained

through launch-vehicle staging, coverage near the region of

staging is considered marginal. At lower altitudes, coverage

was primarily limited in quality by a low-level haze. At

higher altitudes sun, haze, and image reduction caused by the

increasing slant range restricted tracking capability and

affected both the quality and the duration of tracking camera

coverage. Fifteen films were reviewed, including l6-mm and

35-mm film from four fixed cameras and eleven tracking cameras.

Fixed camera coverage with respect to exposure, focus, and film

quality was good, with the exception of one item which faced

into the sun and overexposed and one item which was grainy

because of haze. Two fixed cameras indicated normal lox

boil-off, umbilical disconnect, and umbilical door closure.

The two other fixed cameras showed close-up views of space-

craft and launch-vehicle displacement through lift-off. The

quality of the tracking camera coverage was generally good

with respect to exposure, focus, and tracking but was poor

in quality with respect to color, grain, and resolution.

Five tracking cameras showed launch-vehicle ignition and

lift-off. Ten tracking cameras indicated normal launch-

vehicle staging.



Page 10 - 12

10.5-3 Documentary film.- Documentary coverage used for engineering

evaluation of the mission was provided "by eight l6-mm motion-

picture films and numerous still photographs. Three films

provided tracking coverage of lift-off from different loca-

tions near complex 12 and from a recovery vehicle positioned

on the beach.

Six aerial motion-picture films of the launch sequence were

taken by aircraft in the launch area. Two films provided very

good aerial photographic coverage with the exception of slight

camera vibration and intermittent tracking. The remaining

aerial films provided additional photographic coverage of the

launch sequence but were poor in quality with respect to color,

focus, and resolution. One motion-picture film of the recovery

operation was-available for review. This film provided aerial

and shipboard photographic coverage of the spacecraft descending

on the main parachute, spacecraft landing in the water, para-

rescue personnel being dropped near the spacecraft, activities

of pararescue team with the spacecraft in the water, space-

craft retrieval from the water by the recovery aircraft carrier,

spacecraft preparation for astronaut egress, removal of the

spacecraft hatch, astronaut egress from spacecraft, and re-

trieval of pararescue personnel from water. Photographs of

personnel with the spacecraft below the main deck of the air-

craft carrier were also obtained.

Documentary coverage of the mission by still photography was

very good. Numerous still photographs were available for re-

view, in which prelaunch, launch, flight, recovery, and post-

flight operations were documented. Still photographs of

prelaunch activities included views of astronaut preparation

at Hangar S, insertion of the astronaut into the spacecraft,

and securing for launch. Also included were prelaunch photo-

graphs of the spacecraft alone and mated with the launch vehicle.

Still photographs taken during the flight provided views of the

launch sequence from different locations, flight operations in

Mercury Control Center, and the slow-scan TV pictures received

from the spacecraft. Recovery photographic coverage showed

views of the spacecraft on the main parachute, the spacecraft

and pararescue personnel in the water before retrieval by the

recovery carrier, spacecraft retrieval by the aircraft carrier,

the spacecraft onboard the carrier after pick up, removal of

the hatch, astronaut's egress, the spacecraft in close-ups

after recovery, and astronaut activities after the medical

examination. Engineering still photographs, showing close-up

views of the spacecraft during postflight inspection at Cape

Canaveral, were also available.



Page 10 - 13

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10.6 Postf light Inspection

Spacecraft 20, shown in figure 10.6-1, underwent the normal

postflight inspection and conditioning procedure. A thorough visual

inspection was made of the external and internal areas of the space-craft in the "as received" condition. The immediate postflight-

inspection procedure included removal of the heat shield, landingbag,

and conical and beryllium shingles for inspection of the pressure

bulkhead and internal skin areas. A photographic record was made of

the inspection process.

A desalting wash-down, tank drainage, and flushing procedure, as

applicable, was accomplished; and safeguards against deterioration

were taken. The detailed inspection results of the individual space-

craft structural systems are discussed in the following paragraphs.

10.6.1 General.- The overall condition of the spacecraft structurewas good. The outer pane of the spacecraft window was

broken during actuation of the explosive hatch aboard the

recovery ship. As on previous flights, there were drops

of water on the inside of the outer pane of the spacecraft

window. The exterior of the spacecraft showed the usual

discoloration due to aerodynamic heating. There were

numerous deposits of molten metal on both the conical and

cylindrical portions of the spacecraft exterior. The

deposits were in the areas above each of the three

retropackage umbilicals and one of the spacecraft-adapter

umbilicals, all of which failed to jettison from the

spacecraft. Most of the deposits had the appearance ofsolder spots. There was also evidence of considerable

aerodynamic heating in the area around where the still

attached umbilicals were located.

The coaxial antenna cable also did not separate from the

spacecraft, although it had been severed by the coaxial

cutter located beneath the shingle. The cable evidently

was caught at the hole in the shingle, through which it

passed and broke off,leaving about 6 inches of free cable

remaining.

Hydrogen peroxide was noted to be dripping very slowlyfrom the spacecraft between stringers 15 and l6 on to the

edge of the heat shield. Subsequent inspection showed that

an RCS line beside stringer 15 had corroded through from

the outside and was leaking hydrogen peroxide. Several

RCS lines had "considerable amounts of corrosion on their

outer surfaces; however, all RCS lines were exceptionally

clean and free of corrosion on their inner surfaces.

COMJD:



Page 10 - 15

10.6.2 Structure.- The spacecraft experienced very little

structural damage. Two cracks in the outer skin of the

pressure vessel, one of which was 3^ inches long and the

other 2 inches long, were noted in the area just to theright of the hatch. This damage most probably resulted

from the explosive-hatch actuation which produced a warpin the hatch sill.

10.6.3 Ablation shield.- The ablation shield appeared to be intact.

There were several minor circumferential cracks noted in

ablation laminate, but they were less severe than the ones

noted on the MA-8 heat shield. One of the heat-shield re-

taining lugs was broken off and another was bent. The

ablation-shield bondline under the flotation weights hadsep-

arated in places, but the bolts held the weights in place.

10.6.^4- Landing bag.-The landing bag was slit vertically in six

places. The slits varied in length from 6 to 20 inches.

In addition, there were numerous small tears and punctures

near the top of the bag. All of the landing bag straps

were intact, although they were twisted and kinked.

Twisting and kinking probably occurred during postrecovery

handling.

10.6.5 Recovery compartment.- Wo damage was noted in the recovery

compartment area except that the UHF descent and recovery

antenna was broken loose at the hinge point. This damage

occurred while the spacecraft was being lifted aboard the

recovery ship.

10.6.6 Main pressure bulkhead.- The main pressure bulkhead area

sustained very little damage on landing. The postflight

appearance of the honeycomb structure was not noticeably

changed from that before the flight. The fiberglass shield

had five scratches along the right X-axis and a dent in the

center stiffening ring in the same location. The protector

over the manual system selector valve was bent; and two dents,

or creases, in a bead in the skin next to the valve indicated

that the skin was pushed in during the landing. Also, there

was a surface crack in a spot weld in this area.

10.6.7 Spacecraft interior.- The interior of the spacecraft was

in good condition. The astronaut's couch had been

removed by the recovery forces for access to the area under

the couch and it had been replaced without being bolted

down. Approximately 0.28 pound of liquid was removed from

under the couch by the recovery personnel, and approximately

0.39 pound of liquid was removed from the same area during



Page 10 - 16

postflight inspection at Cape Canaveral. A chemical

analysis of this liquid to determine its source is in

process. There was a considerable amount of paint chips

throughout the interior of the spacecraft. These chips

resulted from initiating the explosive-actuated hatch.

CONFIDE1



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Page 11-1

APPENDIX B

11.0 ACKNCMIiEDGEMENT

The Flight Evaluation Team for the MA-9flight, upon whose

analysis this report is "based, was composed as follows:

3.0 SPACE VEHICLE DESCRIPTION

3-1 Spacecraft Description

C. Vaughn

3.2 Launch Vehicle Description

L. DuGoffA. E. Franklin

14-.0 TRAJECTORY AND MISSION EVENTS

D. Incerto M. Apple

M. Cassetti F. McCreary

E. Hawkins J. Wells

P. McKaskill S. Yates

5.0 SPACECRAFT PERFORMANCE

5.1 'Spacecraft Control System

G. T. Sasseen

R. Buckley

T. Williams

5.2 Lift Support System

F. Samonski J. Billingham, M.D.

D. Hughes D. Hampton

J. Whalen F. Hettinger

5-3 Communications Systems

W. R. Stelges

5. Mechanical and Pyrotechnic Systems

S. T. Beddingfield

FIDENTIAL



11 - 2 COPTICC M T I A L i

5-5 Electrical and Sequential Systems

M. A. Guidry

J. D. Collner

5-6 Instrumentation System

¥. R. Durrett

M. A. Wedding

H. J. Ness

5-7 Heat Protection Systems

J. Pavlosky

5.8 Scientific Experiments

¥. ArmstrongJ. McKee

6.0 LAUNCH VEHICLE PERFORMANCE

L. DuGoff

M. Cassetti

A. E. Franklin

7.0 ASTRONAUT ACTIVITIES

7-1 Aeromedical Analysis

E. P. McCutcheon, M.D.D. D. Catterson, M.D.

R. A. Pollard, M.D.

H. A. Minners, M.D.

R. Hackworth

7-2 Astronaut Performance

J. B. Jones T. ¥. Holloway

J. J. Van Bockel R. B. Benson

R. D. Mercer G. ¥. Harvey

7.3 Pilot's Flight Report

L. G. Cooper, Jr.

8.0 FLIGHT CONTROL AND NETWORK PERFORMANCE

8.1 Flight Control Summary

C. Kraft

E. Kranz

J. Hodge

Documents

Post Launch Memorandum Report for Mercury-Atlas No. 9(MA-9). Part 1 Mission Analysis