First United States Manned Three-pass Orbital Mission Mercury-Atlas 6, Spacecraft 13 Part 1 - Description and Performance Analysis

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TECHN ICAL

MEMORANDUM

>

N A S A T M X - 5 6 3 - I

C U S S I F K iiT t iia

J>eciasslfled by authority of H A S AClassification Change NoticesDated •

IR S T U N I T E D S T A T E S M A N N ED

- P A S S O R B I T A L M I S S I O N

6 , S P A C E C R A F T 13 )

I - D E S C R I P T I O N AND

F O R M A N C E A N A L Y S I %

nby John H. B o y n t o

C e n t e r

o u st o n , Texas

A E R O N A U T I C S A N D S P A C E A D M I N I S T R A T I O N • WASH|NGTO||t C . • M A R C H 1 9 6 4



TECHNICAL M E M O R A N D U M X-563-1

FIRST U N I T E D STATES M A N N E D THREE-PASS ORBITAL MISSION

( M ER C UR Y - A T LA S 6, SPACECRAFT 13)

PART I - DESCRIPTION ANDPERFORMANCE A N A L Y S I S

Edited by John H. Boynton

Manned Spacecraft Center

Houston, Texas

N A T IO N A L A E RO N A U T IC S A N D SPACE A D M I N I S T R A T I O N



FOREWORD

The first United States manned orbital flight constitutes a significant

milestone in a national program of continued space exploration. The success

of this flight was largely dependent on -the realization of objectives estab-

lished for the two manned suborbital missions and the numerous unmanned

development flights which have been completed as a part of Project.Mercury.

General acknowledgement is made of the extensive effort on the part of

the entire Mercury team. This team, consisting of many organizations that

are external to the Manned Spacecraft Center, notably includes the Department

of Defense, the spacecraft contractor and its subcontractors, the NASA.

Goddard Space Flight Center for the Mercury Worldwide Network, the launch

vehicle contractors and their subcontractors, and in general the many orga-

nizations and government agencies which directly or indirectly made possible

the success of this historic flight. ~-~-

The contents of this volume represent the contributions of an assigned

flight evaluation team, comprising system specialists and operations personnel

from throughout the Manned Spacecraft Center, without whose analytical and

documentary efforts a report of this technical completeness would not have

been possible.



CONTENTS

Page

FOREWORD ........ ...................... i

LIST OF TABLES ........................... vii

LIST OF FIGURES ........................ . . ix

SUMMARY. .............................. 1

INTRODUCTION ............................ 2

SPACE-VEHICLE DESCRIPTION ................ ..... 6

SPACECRAFT DESCRIPTION ...................... 6

Spacecraft Control System ................... 6Environmental Control System .................. 7

Communications Systems ............. ........ 10Electrical and Sequential Systems ....... .. ....... 10Electrical power system ................... 10Sequential system ...................... 1.1

Heat Protection System ..................... 11Ablation shield ....................... 11

Afterbody .......................... 11Mechanical and Pyrotechnic Systems ............... 11

Separation devices ...................... 12

Rocket motors ........................ 13Landing system ........................ 1Înternal spacecraft structure .......... ...... 1^

Instrumentation System ..................... 15Spacecraft Modifications .................... 15

LAUNCH-VEHICLE DESCRIPTION .................... j8

Airframe ............................ 3&Propulsion System ....................... 38Guidance System ... ...................... 38Abort Sensing and Implementation System ............ 39Aerodynamic Load Criteria ......... .......... 39

Launch-Vehicle Modifications .................. 39

MISSION OPERATIONS

PRELAUNCH OPERATIONS ....................... Ul

Astronaut Training and Preparation ............... ^1Academics .......................... ^1Static training devices ................... ^2Environmental familiarization ................ ^2Dynamic training devices ................... ^2Egress and survival training ................. ^3

ii



CONTENTS - Continued

Page

Specific mission preparation ^3

Spacecraft Prelaunch Preparation ^3Time utilization ^3Design changes ^

Systems tests ^Simulated flights ^6Launch pad operations (prior to countdown) ^6

Spacecraft History t ^7 .Reaction control system .' 7Environmental control system ^7Communications ^8Electrical system ^8

Launch-Vehicle Prelaunch Preparation ^9Flight Safety Reviews 50First series of reviews 50Second series of reviews 51

LAUNCH OPERATIONS 56

Launch Procedure 56Weather Conditions 57Photographic Coverage 57

FLIGHT-CONTROL OPERATIONS 6l

RECOVERY OPERATIONS 62Recovery Plans • • • • 62Recovery Procedure 62Recovery Aids 63

MISSION PERFORMANCE 68

SPACECRAFT PERFORMANCE 69

Spacecraft Control System 69System description 69Flight description and analysis 69Powered flight and turnaround 69Orbital phase 70Retrofire 71Reentry 71

Reaction control system 72Prelaunch activities 72Flight performance 72Hydrogen peroxide feed-line temperatures in flight 73Postflight inspection 73

Environmental Control System ? * * •

iii



> • • • • • • • • • • • • •» • %

CONTENTS - Continued

Page

System description 7^Countdown 7^Launch phase . . 7^

Orbital phase 74Reentry and postlanding 75Postflight investigation 75

Communications Systems 76Voice system 76Radar system j6Command system j6Recovery system j6

Electrical and Sequential Systems 76

Electrical system 16Sequential system 77Instrumentation 77Telemetry 78Data quality 78Photographic 78Onboard timing : 78Respiration sensor 78Fuel-quantity indicators 79

Heat Protection System 79Mechanical and Pyrotechnic Systems 79

Parachutes 79

Rockets and pyrotechnics 8lExplosive-actuated hatch 8lLanding-shock attenuation system 8l

Postflight Inspection 8lStructure 8lAblation shield 8lHeat-shield-deployment instrumentation 82Landing bag 82

AEROMEDICAL ANALYSIS ."' 108

Clinical Studies 108

Physiological Studies 110Data sources 110Bioinstrumentation . IllPreflight . IllFlight 112Pilot inflight observations 114

ASTRONAUT FLIGHT ACTIVITIES 133

Preflight Training Summary 133Spacecraft checkout activities 133

iv



•••

CONTEHTS - Continued

Page

Training activities ..................... 133

Flight preparedness ..................... 133Chronology of Pilot's Activities During Flight ........ 133Attitude Control .......................

Control systems check ..... ...............The 60° right-yav maneuver .................Three l80° right-yaw maneuvers ...............Use of a constellation as a reference ............Gyro caging .........................Retrofire control .......... ............ 135

Reentry pitch maneuver ................ ... 135

Reentry damping ..... .................. 135

Pilot's use of external reference .............. 135Communication Activities ................ ... 136

Scientific Observations ....... . ............ 136

Celestial observations ................... 136Meteorological observations ....... .......... 137Terrestrial observations .................. 137

Color photographs .................... . . 138Sensation and Orientation During Weightlessness ...... . . 138

General sensations ..................... 138

Orientation ................ ..... .... 138

Personal Equipment .................... . . 139

Daylight color camera .................... 139

Ultra-violet spectrograph .................. 139Photometer ......................... 139

Airglow filter ....................... 139Night adaptation eye patch ................. 139Flight-plan cards ......................Food tube ................. .........Food tablets .............. ..........

PILOT'S FLIGHT REPORT ......... ............. 156

Preparation and Countdown ................... 156Powered Flight ....... ................. 156

Orbital Insertion ....................... 157Orbit . ............................ 158

Thruster problem ..................k. . . . 158

Attitude reference ..................... 158Weightlessness ....................... 159

Color, light, and visibility ................ 160Space particles .......... . ............ 163

Other planned observations ................. 163Reentry ............................ 164

Landing and Recovery ....... ... ....... .... 166

Concluding Remarks ...................... 166



CONTENTS - Concluded

Page

LAUNCH-VEHICLE PERFORMANCE 168Abort Sensing and Implementation System (ASIS) 168

Engine Cutoff 168

Orbit Lifetime . 168

Guidance 168

Aerodynamic Loads 169

TRAJECTORY AND MISSION EVENTS 169

Sequence of Flight Events 169

Flight Trajectory 169

Launch phase 169

Orbital phase 170

Reentry phase 170

MERCURY NETWORK PERFORMANCE 195

Tracking 195

Data Transmission 195

Trajectory Computation 196

Telemetry 197

Air-Ground Voice 197

Command System 198

Ground system 198

Airborne system 198

Ground Communications 199

CONCLUDING REMARKS 22k

REFERENCES 226

vi



LIST OF TABLES

Table Page

I. SPACECRAFT CONTROL SYSTEM REDUNDANCY AND ELECTRICAL

POWER REQUIREMENTS .................... 18

II. SPACECRAFT COMMUNICATIONS AND INSTRUMENTATION SYSTEM .... 19

III. AMR OPTICAL COVERAGE OF LAUNCH AND REENTRY PHASES ...... 58

IV. FUEL CONSUMPTION ...... ................ 83

V. RESULTS .OF POSTFLIGHT EXAMINATION OF THRUST CHAMBERS .... 8

VT. INSTRUMENTED PARAMETERS FOR MA-6

(a) Commutated quantities .............. . 87

("b) Continuous quantities ................ 90(c) Onboard tape recorder track assignments ....... 90

VII. TELEMETRY SIGNAL STRENGTH ................. 91

VIII. AEROMEDICAL EVENTS PRIOR TO LAUNCH ..... ........ 115

IX. XYLOSE ABSORPTION STUDY .................. 116

X. URINE SUMMARY ............. .......... 117

XI. PREFLIGHT AND POSTFLIGHT PHYSICAL EXAMINATIONS OF THE

ASTRONAUT . . ...................... 118

XII. PERIPHERAL BLOOD ...... . . . ............. 119

XIII. BLOOD SUMMARY ... .................... 120

XIV. PLASMA ENZYMES SUMMARY ................... 121

XV. TIME EXPENDED IN ASTRONAUT PRELAUNCH ACTIVITIES ......

XVI. TRAINING SUMMARY FOR PILOT IN CAPE CANAVERAL

PROCEDURES TRAINER ....................

XVII. CONTROL MODE AND ATTITUDE MANEUVERS DURING MA-6 MISSION

XVTII. NUMBER OF COMMUNICATIONS TO AND FROM SPACECRAFT

XIX. FIIM EXPOSURES

vii



LIST OF TABLES - Concluded

Table Page

XX. SEQUENCE OF EVENTS 172

XXI. COMPARISON .OFPLANNED AND ACTUAL TRAJECTORY PARAMETERS ... 17J

XXII. ORBITAL INSERTION CONDITIONS .DISPLAYED AT MCC 200

XXIII. SUMMARY OF LOW-SPEED TRACKING DATA 201

XXIV. SUMMARY OF LANDING-POINT PREDICTIONS BASED ON RADAR

TRACKING 202

XXV. TELEMETRY RECEPTION SUMMARY

(a) First orbital pass 203

(b) Second orbital pass 20k(c) Third orbital pass 205

XXVI. COMMAND HANDOVER SUMMARY . 206

XXVII. COMMAND FUNCTION SUMMARY 207

viii



UST OF FIGURES - -

Figure Page

1. MA-6 astronaut prior to the mission 3

2. MA-6 astronaut in the spacecraft prior to hatch closure . . U

3. Ground track for the MA-6 orbital mission 5

4. Space-vehicle configuration 20

5 » Spacecraft axis system 21

6. The MA-6 Mercury spacecraft and launch-vehicle adapter . . 22

7. Reaction control system

(a) System A 23(b) System B 2U

8. Environmental control system 25

:9- Location of communication systems 26

10. Spacecraft antenna schematic diagram . . 27

11. Voice system schematic diagram 28

12. Sequence of major events 29

13- Master sequential diagram

(a) Launch and orbit 30(b) Retrograde and reentry 31(c) Landing and recovery 32

lU. Heat protection system 33

15. Rocket motor ignition circuitry 3^

16. Instrumentation system diagram 35

17. Instrumentation sensor locations 36

18. Instrument panel 37

19. Time analysis of the spacecraft prelaunch operations ... 52

20. Launch complex testing of the MA-6 spacecraft ........ 53

ix



LIST OF FIGURES - Continued

Figure Page

21. Launch complex modifications 5^

22. Emergency egress tower 55

2J. Wind profile at the launch site 59

2k. AMR engineering sequential tracking camera coverage . . . . 60

25- Recovery areas and ship locations 64

26. Contingency recovery support forces 65

27. Details of primary landing area 66

28. Spacecraft prior to"retrieval 67

29. Example of the l80° yaw maneuver 92

30. Fuel consumption during reentry (within 0.5 lb) 93

31. Postflight photograph showing 1-pound yaw thruster hardware

(a) Fuel metering orifice 94(b) Fuel distribution screen . 94

32. Variation of suit-inlet, cabin, and inverter temperatureswith time

(a) Suit-inlet and cabin temperature 95(b) Inverter temperatures 95

33- Variation of primary and secondary oxygen pressures withtime

(a) Prelaunch 96

(b) During flight 96

34. Reentry-heating time history for MA.-5

(a) Cylindrical section 97

(b) Conical section, no. 1 98

(c) Conical section, no. 2 99

(d) Conical section, no. 3 100

(e) Conical section, no. 4 101

(f) Conical section, no. 5 102

(g) Conical section, no. 6 103




Figure Page

(h) Conical section, no. 7 104

35. Postflight photograph of MA-6, left-hand limitsvitch shaft 105

36. Postflight photograph of spacecraft 106

37. Postflight photograph of a"blation shield 107

38. Respiration rate, pulse rate, body temperature, suit-inlettemperature, and "blood pressure measured duringcountdown

(a) Countdown, 06:00 to08:00 e.s.t. . . . . . . ... . 122Ob) Countdown, 08:00 to 09:47 e.s.t. (lift-off) ..... 123

39' Sample of MA.-6 blockhouse countdown recorded at the timeof astronaut insertion (T-220 min). Recorder speed,25 mm/sec 124

40. Record of MA-6blockhouse preflight blood pressure takenbefore lift-off (T-50 sec). Recorder speed, 25 mm/sec . . 125

41. Respiration rate, pulse rate, body temperature, and

blood pressure during flight

(a) Ground elapsed time, 00:00 to 02:30 126(b) Ground elapsed time, 02:30 to 05:00 127

42. Sample of physiological record received at Bermuda trackingstation during powered phase of flight at approximately4 minutes after lift-off. (Recorder speed, 25 mm/sec) . . 128

^3. Physiological data after 2 hours and 50 minutes ofweightlessness, including inflight blood-pressure

trace (value, 134/64 mm Hg).

(a) Sample of onboard record (recorder speed, 10 mm/sec). 129(b) Sample of telemetered data received at Hawaii .... 130

44. Sample of onboard record of physiological data at drogueparachute deployment, approximately 4 hours 49 minutesafter lift-off. (Recorder speed, 25 mm/sec) . . . . . . . 131

45. Inflight exercise device . . . . . . . . . . 132

xi




Figure Page

k6. Chronology of pilot activities

(a) First orbital pass .(b) Second orbital pass(c) Third orbital pass

Vf. One 60° right-yav maneuver using the periscope andfly-by-wire .......... ............. 150

8. Three l80° right -yaw maneuvers using the window

reference and manual proportional or fly-by-wire .... 150

49. Control of vehicle attitudes using the constellation Orionthrough the window as a night reference system andfly-by-wire control mode .... ............ 151

50. Reentry maneuver using the manual proportional controlmode and instrument reference ... ........... 151

51. Onboard, record of the two high oscillation periods

during reentry;

(a) Maximum reentry oscillation periods during reentry . 152

(b) Second oscillation period ............. ' 153

52. Personal equipment

(a) Color camera(b) Ultra-violet spectrograph(c) Photometer(d) Airglow filter ................... 15U

, (e) Night adaption eye patch .............. 155(f) Food tubes ..................... 155(g) Food tablet dispenser . ..... ........ . 155

53. Comparison of MA .-5and MA .-6 spacecraft pitch ratesduring launch ...................... 175

5 . Inertial velocity and flight-path angle in the regionof cutoff using launch-vehicle guidance-system data

(a) Inertial velocity ................. 176(b) Inertial flight-path angle ...... ....... 177

55. Inertial velocity and flight-path angle in the region ofcutoff using I. P. 7090 data.

xii



• •«

• •«

• »i


Figure Page

(a) Inertial velocity . . . . . . . . . . . . 178

(b) Inertial flight-path angle 179

56. Inertial flight-path angle plotted against inertialvelocity in the region of cutoff l80

57- Calculated values for oq. for the MA-6 launch l8l

58. Altitude plotted against longitude profile . . 182

59« Time histories of trajectory parameters for MA-6mission launch phase .

(a) Altitude and range plotted against time l8j(b) Inertial velocity and flight-path angle plotted

against time . . . . . . . . . . l8U(c) Earth-fixed velocity and flight-path angle plotted

against time 185(d) Dynamic pressure and Mach number plotted against

time . . 186(e) Longitudinal acceleration along the spacecraft . -

.Z-axis plotted against time . . . . . . . . . . . . . 187

60. Time histories of trajectory parameters for MA-6

mission orbit phase

(a) Latitude, longitude, and altitude plotted againsttime 188

(b) Inertial velocity and flight-path angle plottedagainst time . 189

61. Time histories of trajectory parameters for MA-6mission reentry phase

(a) Latitude, longitude, and altitude plotted against

time 190(b) Inertial velocity and flight-path angle plottedagainst time 191

(c) Earth-fixed velocity and flight-path angle plottedagainst time 192

xiii



UST OFFIGURES - Continued

Figure

(d) Dynamic pressure and Mach number plotted against

time 193(e) Longitudinal deceleration along spacecraft Z-axis

plotted against time

62. S-"band radar coverageb


(b) Second orbital pass 209(c) Third orbital pass 210

63. C-band radar coverage


(b) Second orbital pass 212

(c) Third orbital pass 213 ~

64. Telemetry reception coverage

'a) First orbital pass 2lkb) Second orbital pass 215c) Third orbital pass 2l6

65. Fuel quantity, automatic and manual systems 217

66. HF and UHF voice coverage


(b) Second orbital pass 219(c) Third orbital pass 220

67. Recorder command functions during ionization blackout . 221

xiv



NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

TECHNICAL MEMORANDUM X-56}

FIRST UNITED STATES MANNED THREE-PASS-ORBITAL MISSION

( M E R C U R Y - A T L A S 6 , S P A C E C R A F T 1 3 )

Part I - D E S C R I P T I O N A N D P E R F O R M A N C E A N A L Y S I S *

Edited by John H. Boynton

SUMMARY

The Mercury-Atlas mission 6 was.the first United States manned orbital

flight. A detailed discussion of the mission, including the preflight

operations, and a comprehensive postlaunch evaluation are presented. Only

data vhich significantly amplify the context are included.

All prescribed mission objectives were successfully accomplished and a

comparison of the planned and actual trajectory data indicates that pertinent

mission parameters nearly coincide with expected values. The spacecraft,

launch vehicle, and Mercury Worldwide Network functioned satisfactorily

throughout the mission.

A number of minor discrepancies occurred but they did not compromise

the success of the mission. Early in the flight, a telemetry signal was

received by ground stations which indicated the heat-shield release mechanism

had been actuated. A postflight investigation revealed that a faulty limit

switch had caused the misleading telemetry signal. During the second orbital

pass, the periodic loss of automatic spacecraft stabilization because of a

failure of the 1-pound yaw thrusters required the astronaut to control the

spacecraft manually. Oscillations of the spacecraft in pitch and yaw during

reentry increased to unexpected levels and were subsequently eliminated

through deployment of the drogue stabilization parachute. Drogue parachute

deployment did, however, occur earlier than had been planned. Despite these

and other minor anomalies, the mission was completed successfully, and recoverywas effected in the prescribed area within 20 minutes after landing.

The astronaut satisfactorily monitored and controlled the operation of

spacecraft systems, performed planned attitude maneuvers, and observed

terrestrial phenomena. Throughout the flight, the astronaut's physiological

and psychological responses to the orbital space environment were within normal

ranges, and his health following the flight has remained excellent.

Title, Unclassified.



> »• « •• ft

HWRODUCTION

The first manned orbital flight of the Mercury program was successfullyperformed on February 20, 1962. Astronaut John H. Glenn, Jr., shown in

figures 1 and 2, was the assigned pilot for this mission. Figure 1 depictsthe astronaut in his full-pressure suit with the portable cooling unit. Aswell as being the third orbital flight of a Mercury spacecraft, Mercury-Atlasmission 6 (MA-6)marked the sixth of a series of flights utilizing specificationMercury spacecraft and Atlas launch vehicles. The MA-6space vehicle waslaunched from the Cape Canaveral Missile Test Annex in Florida.

The MA-6mission was planned for three orbital passes, with the groundtrack illustrated in figure 3, and was the culmination of a program to developthe Mercury spacecraft for manned orbital flight. The objectives of theflight were to evaluate the performance of the man-spacecraft system in athree-pass mission, to evaluate the effects of space flight on the astronaut,

and to obtain the astronaut's evaluation of the operational suitability ofthe spacecraft and supporting systems for manned orbital missions.

All data telemetered and recorded during the flight have been thoroughlyanalyzed by system specialists, and this report presents these results andtheir analyses. Brief descriptions of the spacecraft, the launch vehicle,and the operations necessary to the mission precede the performance analysisand supporting data. All significant events of the MA-6 mission, beginningwith delivery of the spacecraft to the launch site and concluding with therecovery and postflight examinations, are documented.

Lift-off for the MA-6 mission occurred at 9 hours, ^7 minutes, and39 seconds a.m. e.s.t. All times throughout this report are given as groundelapsed time (g.e.t.) from lift-off, unless otherwise noted.

Although the graphical information presented in this part of the reportsufficiently supports the text, part II of this report contains a completepresentation, without analysis, of all MA-6time-history data.



• • • <• • •

Figure 1. - MA-6 astronaut prior to the MA-6 mission.



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SPACE-VEHICLE DESCRIPTION

The lift-off configuration for Mercury-Atlas mission 6 consists of theMercury spacecraft and the Atlas D launch vehicle. (See fig. 4.) A general

description of these vehicles is presented subsequently, along with a listingof significant spacecraft modifications since the previous Mercury-Atlasmission (MA-5), which was an unmanned orbital flight with a chimpanzee on-board. The descriptions in this paper are presented merely for the purposeof familiarizing the reader with the systems whose performance during themission is discussed in subsequent sections; however, a more thorough de-scription of the spacecraft is presented in references 1 to 3 and papers2 and 3 of reference k.

SPACECRAFT DESCRIPTION

The Mercury spacecraft is designed to provide a safe and habitable

environment for the pilot in space, as well as protection during the criti-

cal flight phases of launch and reentry. The spacecraft also serves as an

orbiting laboratory where the pilot can conduct limited scientific experi-

ments concerning the space environment. The axis system for the Mercury

spacecraft is depicted in figure 5- The MA.-6 spacecraft (no. 13), shown

just prior to launch in figure 6, was nearly identical to the spacecraft

utilized for the previous orbital missions. The many systems which the

spacecraft comprises may be generally grouped into those of spacecraft

control, environmental control, communications, mechanical and pyrotechnic,

electrical and sequential, heat protection, and the onboard instrumentation.

Spacecraft Control System

The spacecraft control system provides the capability to achieve and

maintain closely precise attitude during the orbital, retrofire, and reentry

phases of the flight. Because the retrofire maneuver is so critical to the

mission, the control system has been designed so that it can perform its

function in the event of multiple system malfunctions.

Table I lists the four control arrangements that are available in the

spacecraft. For the reaction control system (RCS), there are two completely

independent fuel-supply, plumbing, and thruster systems, and the locations

of these components are indicated in figure 7- Each uses 90-percent hydrogen

peroxide to provide selected impulse as desired. There are two means of con-

trolling the outputs of each of these systems; that is, on system A the astro-

naut has a choice of using either the automatic stabilization and control

system (ASCS) or the fly-by-wire (FEW) system. The ASCS is automatic to the

extent that it can provide the necessary attitude control, including fixed

orbital precession to maintain a constant angle with respect to the local

vertical, throughout a complete mission without any action on the part of

the astronaut. The ASCS derives its attitude reference from the spacecraft



gyros, which are in turn slaved to the horizon scanners to eliminate gyroprecession errors. The FEW system is operated by movement of the astronaut'scontrol stick which is linked electrically to the solenoid control valvesof system A.

For system B, the astronaut has the choice of using either the manual

proportional (MP) or the rate stabilization control system (RSCS) modes,both of which are operated through the astronaut's control stick. In theMP system, linkages transmit the control stick movement to proportionalcontrol valves which regulate the flow of fuel to the thrusters. The RSCSuses the combination of control stick inputs and the computing componentsof the automatic control system to provide rate control.

The desired control mode can be easily selected by positioning theproper switches and valves mounted on the instrument panel. Certain com-binations of these control modes can be selected to operate simultaneously,such as ASCS and MP, or FEW and MP, in order to provide double authoritycontrol, so that even with certain malfunctions in each mode adequate control

can be maintained.

The thruster impulse is directed by the four basic control modes through18 individual thrusters - 12 on system A (automatic-fuel system) and 6 onsystem B (manual-fuel system). Schematic diagrams of system A and B arepresented in figure 7- Metered quantities of hydrogen peroxide are decomposedin silver-plated catalyst beds in each of the thruster chambers to providethe desired impulse. Twelve of the thrusters used on the Mercury spacecraftare sized to provide adequate control during the critical retromaneuver.These RCS thruster ratings are as follows:

Axis

Pitch

Yaw

Roll

System A, Ib

24

2k

6

System B, Ib

4 to 24

4 to 24

1 to 6

The remaining six thrusters are in system A and each has a thrust rating of

1 pound. Under orbital conditions, these thrusters provide fine attitudecontrol as required.

Environmental Control System

The Mercury environmental control system (ECS) provides a livableenvironment for the astronaut in which total pressure, gaseous composition,and temperature are maintained and a breathing oxygen supply is provided.To meet these requirements, a closed-type environmental control system wasdeveloped.



•I

••

The environmental control system shown in figure 8 is located in thelower portion of the spacecraft under the astronaut support couch. Theastronaut is clothed in a full-pressure suit to provide protection in theevent of a cabin decompression.

The pressures in the cabin and pressure suit are maintained at 5.1 psiain normal flight with a 100-percent oxygen atmosphere. The system is designed

to control automatically the environmental conditions within the suit andcabin throughout the flight. Manual controls are provided to enable systemoperation in the event of an automatic system malfunction. The ECS can beconsidered as two subsystems: the pressure-suit control system and the cabinsystem. Both of these ystems operate simultaneously from common coolant-waterand electrical supplies. The coolant water is stored in a tank with apressurized bladder system to facilitate the flow of water into the heatexchangers. Electrical power is supplied from an onboard power supply.

Oxygen is supplied at an initial pressure of 7>500 psi from two sphericalsteel tanks.

The pressure-suit control system provides breathing oxygen, maintainssuit pressurization, removes metabolic products, and maintains, throughpositive ventilation, gas temperatures.

As shown in figure 8, the pressure suit is attached to the system bytwo connections, the gas inlet connection at the waj.st and the gas exhaustat the helmet. Oxygen is forced into the suit distribution ducts, carriedto the body extremities, and permitted to flow freely back over the body tofacilitate body cooling. The oxygen then passes into the helmet where themetabolic oxygen, carbon dioxide, and water vapors are exchanged. The gasmixture leaves the suit and passes through a debris trap where particulatematter is removed. Next, the gas is scrubbed of odors and carbon dioxide

in a chemical canister of activated charcoal and lithium hydroxide. Thegas then is cooled by a water-evaporative type of heat exchanger whichutilizes the vacuum of space to cause the coolant water to boil at approx-imately 35° F- The heat-exchanger exit gas temperature is regulated throughmanual control of the coolant-water flow valve. The resulting steam isexhausted overboard.

The steam exit temperature on the overboard duct is monitored by athermal switch which actuates a warning light when the duct temperaturedrops below 47° F.

The light is on the instrument panel and provides a visual indicationof excessive water flow into the heat exchanger. Proper monitoring of thelight and correction of the water flow rate will prevent the heat exchangerfrom freezing.

In the gas side of the heat exchanger, water vapors picked up in the suitare condensed into water droplets and are carried by the gas flow into amechanical water-separation device. The water separator is a sponge devicewhich is squeezed periodically to remove the metabolic water from the system.This water is collected in a small tank. The constant flow rate of the

atmosphere is maintained by a compressor.

8



In the MA-6 spacecraft, a constant bleed orifice was provided between

the oxygen supply and the pressure-suit control system. This constant oxygen

flow was in excess of astronaut metabolic needs and thus provided a continuous

flushing of the pressure suit to insure adequate oxygen partial pressure. In

normal operation, suit pressure levels were maintained slightly above cabin

pressure by metering this excess oxygen flow through an exhaust port in the

demand regulator. In the event of a cabin decompression, the demand regu-lator would have automatically established a referenced pressure of h.6 psia

for the exhaust port of the regulator, and suit pressure would have been

maintained at this level. The addition of the oxygen bleed orifice was the

major ECS change prior to the MA-6 flight.

An additional mode of operation is provided by the emergency rate valve.

This valve provides an open-type pressure-suit operation similar to aircraft

pressure-suit systems. A fixed flow of oxygen is directed through the suit

for ventilation and metabolic needs. The remainder is vented into the cabin.

This system is used in the event the pressure-suit control system fails and

also during the final stages of descent. The other components of the suit

system are closed off during this mode of operation.

Oxygen is supplied from two tanks, each containing oxygen sufficient for

more than 19 hours. The tanks are equipped with pressure transducers to pro-

vide data on the supply pressure and are connected in such a way that depletion

of the primary supply automatically provides for supply from the secondary

bottle.

The cabin circuit of the ECS controls cabin pressure and temperature.

A cabin relief valve controls the upper limit of cabin pressure. This valve

permits cabin pressure to decrease with ambient pressure during launch until

a level of 5-5 psi has been reached. This valve then seals the cabin at5-5 psia. In addition, a manual decompression feature is incorporated in

this valve to permit the astronaut to reduce the cabin pressure rapidly if

a fire or accumulation of toxic gases occurs.

A cabin-pressure regulator meters oxygen into the cabin to maintain the

lower limit of pressurization at 5.1 psia. A manual recompression feature

is incorporated in the regulator for cabin repressurization after the cabin

has been decompressed. Cabin temperature is maintained by a fan and heat

exchanger of the same type as that described for the pressure-suit system.

Postlanding ventilation is provided through a snorkel system. Following

reentry, at an altitude of 20,000 feet, the snorkel valves open and ambient air

is drawn by the suit compressor through the inlet valve. The gas ventilates

the suit and is released from the spacecraft through the outlet valve.

The astronaut support couch, which is constructed as a glass fiber

laminate, provides normal in-flight support and protects the pilot during

peak acceleration periods. Each couch is individually tailored to the flight

astronaut and is supported on the large pressure bulkhead by crushable aluminum

honeycomb to absorb landing loads. For a description of the pressure suit and

the astronaut's personal survival equipment, see paper 3 of reference k.



•••••

Communications Systems

The spacecraft communications systems consist of radar beacons and voice,command, and recovery links. Each system has main and backup (or parallel)equipment for redundancy, with manual selection of the desired system providedthrough switches mounted on the instrument panel. Table II is a list of

these systems., and figure 9 shows the physical location of the communicationsequipment in the spacecraft. A simplified schematic diagram showing thevarious communications systems and their respective antenna systems is shown

in figure 10.

Three separate voice systems were available to the astronaut - HP, UHFmain, and UHF backup - as shown in figure 11. A redundant ground-to-airvoice link is also available through the command receiver channel. Anadditional air-to-ground communication link is available to the astronautby keying the high-frequency telemeter carrier. It should be noted that allof these links use the main bicone antenna through the use of a multiplexer.

The radar system consists of C- and S-band beacons onboard the spacecraft.Either or both beacons may be interrogated when within range of the appropriateground station.

The command system provides a means of commanding an abort, retrofire,spacecraft-clock change, or instrumentation calibration from the ground.The onboard command system consists of two identical receivers and decoders,each capable of performing any required functions.

The recovery system comprises an HF transceiver (l watt), one recoverypackage containing the CW SEASAVE beacon (l watt), a pulse modulated SARAH

beacon (7-5 watts), and a pulse modulated Super-SARAH beacon (91 watts).The antenna systems used by the recovery system are shown in figure 10.

Electrical and Sequential Systems

Electrical power system.- The electrical power system comprises threemain batteries, two standby batteries, and one isolated battery. The firsttype has 3,000 watt-hour capacity, while the last two types have a1,500 watt-hour rating; all three power sources are silver-zinc batteries.The standby batteries have taps which power the 6-, 12-, and l8-volt busses.Nominal discharge rate for these batteries is k . ^ > amperes, although each

battery is capable of discharging up to ^2 amperes for brief periods.

The two standby batteries are operated simultaneously and in parallel with themain batteries throughout the flight. The isolated battery is held in reserve.

All electrical power sources may be manually switch-operated by the astronaut.

The three inverters installed in the spacecraft, one for the ASCS, onefor the ECS fans, and the third as a standby unit, provide 115-volt, UOO-cycle,single-phase alternating current. One of the two primary inverters has a150 volt-ampere rating, and the other has a 250 volt-ampere rating. Thestandby inverter also has a 250 volt-ampere rating. All power circuits,except the manual standby inverter circuit, are properly fused, and the

10



majority of these circuits may be enabled by the astronaut.

Sequential system.- The spacecraft sequential system automaticallyinitiates flight events and sequences, with a manual override capabilityprovided the astronaut and a backup provision included for certain criticalevents through an air-ground radio link. Figure 12 shows the sequence of

major events for the MA-6 mission, and figure 13 displays the master sequentialsystem for the spacecraft. The astronaut does not have the capability toinitiate sequences during powered flight regarding the launch vehicle, withthe exception of engine cutoff, which is accomplished through the manualabort circuit.

Heat Protection System

During flight through the atmosphere at launch and reentry, the highvelocities generate excessive heat from which the crew and equipment mustbe protected. The spacecraft must also be capable of withstanding the heatpulse associated with the ignition of the launch escape rocket. To provide

this protection, the spacecraft afterbody is composed of a double-wallstructure with thermal insulation between the two walls, and the forebodyor blunt end of the spacecraft is fitted with an ablation-type heat shield.

Ablation shield.- The material for the ablation shield is a mixture ofglass fibers and resin in such proportions that, while the resin ablates,the fibers provide the necessary structural integrity for the shield. Theenergy required in the fusion and vaporization processes of the resin isextracted from the boundary layer, which in turn reduces the temperature ofthe gaseous flow about the heat shield and thereby keeps the interior tem-perature of the spacecraft at a tolerable level. The heat shield has been

designed to withstand a much greater heat flux than that expected for anormal orbital reentry, where less than half of the heat-shield materialis boiled away.

Afterbody.- The spacecraft afterbody, including the conical and cylin-drical sections and the antenna canister, protects the interior of thespacecraft from excessive heating through a construction of high-temperaturemetallic shingles. The conical afterbody shingles and the inner walls of thepressure vessel form a double-wall structure, with insulation material placedbetween the walls. This construction is exhibited in figure lU. The cylin-drical section is fitted with beryllium shingles which employ the heat-sinkprinciple for heating protection. These shingles are mounted such that they

are free to expand and contract without affecting the major load-carryingportions of the spacecraft structure. The conical section and antennacanister utilize Rene' ij-1, which is a high-temperature alloy. The spacecraftwindow, constructed of four thicknesses of high-temperature glass, is designedto withstand the temperature levels during both exit and reentry.

Mechanical and lyrotechnic Systems

The mechanical and pyrotechnic system group consists of the separationdevices, the rocket motors, the landing system, and the internal spacecraft

11



• " • * M• • •• fr\

structure. The landing system includes the drogue stabilization, main andreserve parachutes, and a landing shock -attenuation system.

Separation devices.- Separation devices generally use explosive chargesto effect separation or disconnection of spacecraft components. The majorseparation points are at the interfaces between the spacecraft and launch

vehicle, between the spacecraft and the escape tower, at the heat shield,and around the spacecraft hatch.

The spacecraft -adapter clamp ring attaches the spacecraft to thelaunch vehicle adapter. The clamp ring secures the spacecraft to the adapterthroughout the powered phase of the flight until spacecraft separation bymeans of explosive bolts. The clamp ring consists of shaped segments whichmate with the fiber-glass attach ring of the spacecraft forebody and withthe upper support ring of the adapter. Three explosive bolts, with dualignition provisions, connect the three clamp-ring segments in tension. Ametal striker bracket is bolted every 120

eto the inside of the clamp ring.

When the clamp ring is installed, the striker brackets depress the spacecraft

ring separation -sens ing switches located in the outer periphery of thespacecraft forebody. The exterior of the clamp ring is covered with a heatshield that protects the explosive bolts from excessive heating. The seg-mented fairing assemblies are bolted to the adapter clamp ring. Six cablestraps protect the spacecraft and the launch vehicle from damage by retainingthe clamp ring to the adapter when the explosive bolts are ignited.

The escape-tower-spacecraft clamp ring consists of three segments thatclamp the escape -tower attach ring to the recovery compartment (cylindricalsection) flange. Three explosive bolts, with dual ignition capability,

connect the ring segments in tension. The escape -tower-spacecraft clamp

ring is basically the same in design as the spacecraft -adapter clamp ring,but it is considerably smaller in size. The clamp ring retains the escapetower to the spacecraft until the clamp-ring explosive bolts are ignited.Only one bolt must ignite to effect separation. An aerodynamic -stabilitywedge attached to the clamp ring aids in stabilizing the spacecraft duringatmospheric aborts . Six cable straps , bolted to the escape tower and theclamp-ring stability wedge, prevent structural damage to the spacecraft byretaining the clamp-ring segments of the escape tower when the explosivebolts are ignited.

The ignition of two squib valves allows nitrogen pressure of 3>000to operate the two heat-shield release mechanism actuators. This action

releases the heat shield from the spacecraft. As the heat shield is released,two limit switches sense heat-shield motion and close to energize the landing-bag extension-signal relay. When the actuator piston reaches full travel,it is locked by a spring-loaded pin. The landing-bag circuit is deenergizedwhile the spacecraft is in orbit. Placing the landing -bag switch in the"automatic" position during spacecraft reentry permits normal operation ofthe entire landing system.

An entrance hatch is located on the right side of the conical sectionas viewed from the crew member station. A primer cord placed between hatch

12



I •>••> • •

and spacecraft sill, is provided to release the hatch quickly and enable theastronaut to egress rapidly. The primer cord igniter, located in a cornerof the hatch, is linked to an internal release control initiator. Prior tolaunch, the hatch is bolted and sealed into position with bolts, and twocorrugated shingles are installed over the hatch. The bolts are insertedinto threaded holes in the spacecraft sill. A magnesium gasket, with inlaid

rubber, forms the hatch seal when the hatch is bolted into position. These twoshingles are attached to the hatch stringers, but not to the spacecraftshingles. Following landing, the astronaut may remove the cap from theinitiator and the safety pin from the initiator plunger. By depressing theinitiator plunger, two spring-loaded firing pins strike the percussion capsand detonate the explosive charge which separates the hatch from the space-craft. An exterior hatch-release control is also provided to enable groundpersonnel to release the hatch.

Rocket motors.- The rocket motor assemblies used in the Mercury space-craft employ solid-propellant fuel and are listed in the following table

with their nominal performance characteristics.

Rocket motor

Escape

Tower jettison

Posigrade

Retrograde

Numberof

motors

1

1

3

3

Thrusteach,Ib

52,000

800

^4-00

1,000

Approximate burning

time for each,

sec

1

1.5

l

10

The escape rocket motor is mounted at the top of the escape tower andincorporates three exit nozzles which are canted 19° to direct the exhaustgases away from the spacecraft. The system is optically alined prior to

launch.

The tower-jettison rocket motor also has a three-nozzle assembly, witheach nozzle canted 30°, and is attached to the bottom of the escaperocket-motor case. Its function is primarily to jettison the expended escape

motor case and tower following an abort during launch.

The three posigrade and three retrograde rocket motors are assembledin a package which is located at the center of the heat shield and held to the

spacecraft at the edge of the heat shield by three straps. The posigraderockets are ignited simultaneously to separate the spacecraft from thelaunch vehicle. The retrograde rocket motors are sequentially ignited(5-second delays between motor ignitions) and provide the velocity decrementnecessary to effect reentry.



• • • •

• • ••

All rocket motors have dual ignition systems with independent electricalpower sources. In addition, each ignition system has dual squibs to insureignition. As a typical example, the rocket motor ignition circuitry for theretrograde system is shown in figure 15-

Landing system.- The landing system includes the drogue (stabilization)

parachute, the main and reserve parachutes, and the landing shock-attenuationsystem (landing bag). The latter system attenuates the force of landing byproviding a cushion of air through the deployment of the landing bag andheat-shield structure, which is supported by straps and cables. The landingsystem is armed when the escape tower is jettisoned during exit flight;however, it is not actuated until the spacecraft returns to the denser partof the atmosphere, normally at an altitude of about 21,000 feet. The systemcan be actuated manually by the pilot or automatically.

A dual barostat, which consists of two pressure-sensing devices in parallel,is used to initiate the landing-system deployment sequence automatically. In

this sequence, the drogue parachute is deployed at an altitude of approximately21,000 feet to decelerate and stabilize the spacecraft. At about 10,000 feet,the antenna section and drogue parachute are jettisoned by signals fromanother dual barostat. When the antenna canister is released, it deploysthe main parachute simultaneously into a reefed condition at about 12 percentof the maximum diameter for approximately h seconds to minimize the openingshock. The reserve parachute may be deployed manually by the astronaut inthe event that the main parachute has malfunctioned. Immediately followingmain parachute deployment, the landing bag is extended to attenuate landingloads. After landing, the main parachute is disconnected automatically, andthe reserve parachute is ejected from the recovery compartment.

The drogue parachute has a ribbon-type canopy with a 6-foot diameter,and it is deployed on a 30-foot-long riser. The main and reserve parachutesare essentially identical, with diameters of 63 feet, and either one willprovide a sinking velocity of 30 fps. The canopies of these parachutes are ofthe ringsail type and are constructed of a special nylon fabric.

The landing bag is a rubberized cloth assembly about U feet in length.Prior to deployment, the heat shield is securely attached to the spacecraftby a multiple-contact, mechanical latching system, and the landing bag isfolded and stored in the space between the heat shield and the large pressurebulkhead of the spacecraft. After release, the landing bag reduces landingaccelerations, and after landing, the bag fills with water to provide dynamicstability for the spacecraft.

Internal spacecraft structure.- For convenience, any spacecraft structuralassemblies not a part of the heat protection system or directly related toan operating spacecraft system are classified as internal spacecraft structure.This structure notably includes the large and small pressure bulkheads, con-structed of titanium and aluminum, respectively, as well as internal bracketry.The pressure vessel which houses the astronaut and spacecraft equipment iscomposed of these two bulkheads and the internal skin of the conical sectionwall.



••• ^ ^ ^ r i . . -. . .. . •• •• * » « B B « • • ••• •• • '••• '., ... .« .-• I • • • • • • • • ••• ••

Instrumentation System

The instrumentation system senses information pertinent to over 100 itemsthroughout the spacecraft. The pilot's biological parameters, consisting ofelectrocardiogram (ECG) traces, respiration rate and depth, body temperature,and blood pressure, are of primary concern to flight-control personnel. In

addition, many operational aspects of the spacecraft systems are monitored.These aspects include significant sequential events, control system operation,attitudes and angular rates, electrical parameters, ECS pressures and temper-atures, accelerations along all three axes, and temperatures of systems through-out the spacecraft. These quantities are both transmitted to Mercury Networkstations and recorded on board the spacecraft. Most of the transmitteddata are commutated through 90 segments, while the remaining data, primarilyaeromedical information, are transmitted continuously. A schematic diagramfor the instrumentation system is shown in figure 16, and the locations ofprimary sensors throughout the spacecraft are depicted in figure 17- Alsoincluded in the instrumentation system is a 1.6-vsm. motion picture camerawhich photographs the astronaut from the instrument panel. Since the filmcapacity is limited, the frame speed is dependent on the mission phase. Ahigher frame speed (360 frames per minute) is programed for the more criticalflight phases, such as launch and reentry, and a speed of 5 frames perminute is provided during the orbital phase, with short periodic "bursts ofthe higher speed. Many of the instrumented parameters have a direct readout

on the spacecraft instrument panel (fig. 18) for monitoring by the astronaut.Periodic reports of displayed quantities to ground personnel offer importantverification of the telemetered data during the flight.

Spacecraft Modifications

Although essentially identical to spacecraft used in previous Mercurymissions, the MA-6 spacecraft differs from that used in the MA-5 mission, theorbital flight of a chimpanzee, in the following minor aspects:

1. An astronaut's couch was installed.

2. A personal-equipment container was installed.

3 - Filters were provided for the window.

k. An indicator was added to the instrument panel which displayedthe temperature of the suit-circuit steam vent.

5. The suit-circuit oxygen supply incorporated a constant-bleedorifice (750 cc/min).

6. The suit-inlet snorkel door was located on the conical after-body.

7. Cooling plates of a new design were installed under the primaryinverters, and stainless steel check valves were installed inthe coolant line in lieu of aluminum valves.

15



8. Improved heat-conduction paths and heat sinks were installed

near the thrusters for temperature control of the roll-thrusterfuel lines.

9 - Indicating lights were added to the instrument panel to show

which inverters were operative.

10. For the electrical inverter circuit, manual control switcheswithout fuses were incorporated, since fuse protection is providedelsewhere.

11. The maximum-altitude sensor was provided with a separatebattery.

12. One-ohm fuses were installed, in all squib circuits.

IJ. An integrating accelerometer was installed.

Ik. An Ultra-SARAH beacon was incorporated in the astronaut'ssurvival kit.

15- A reserve parachute was installed.

16. A switch was installed to allow manual override of heat-shielddeployment.

17. The escape-tower legs were of heavier construction.

18. A manually actuated blood-pressure measuring system wasincorporated.

16



••• •

•• •• • •

The weight and balance parameters for spacecraft 13 are shown in thefollowing table:

Parameter

Weight Ib

Center-of -gravitystation, measuredfrom an arbitrarystation along the

Z-axis below thespacecraft on thelaunch vehicle, in.

z

x

Moments of inertia,

slug -ft

Iz

IX

Iy

Lift-off

k,265 26

167.96

0.31

-0 08

384 0776l

7 767 51 )1w 1 • S

MissioiBeginning offirst orbital

pass

2 986 78

121. 18

-0 Ok

o 0 7

28l6

6216

f i ? Q - * >ucy, y

i phase

Beginning of

reentry

2 698 98

IPk £p

-0 07

o m

P 7 i n

cjLL. £

SS9 PJJ'-'C-

Flotation

(a)

P kpl 7QC^ ,-TC-L. ( y

1 T Q 7k-L-L . ft-

_n - Z . - Z .**"'JJO -t f.

OÂ ld.j\j, J.

XR^ kJJJ>^

^RQ Rj 5 p y . p

Values represent the spacecraft in the dry condition, landing bag andantenna deployed.

IT



TABLE I.- SPACECRAFT CONTROL SYSTEM REDUNDANCY

AND ELECTRICAL POWER REQUIREMENTS

Control systemmodes

ASCS

FEW

MP

RSCS

Corresponding fuelsystem (fuel supply,plumbing, and thrusters)

A

A

B

B

Electrical poverrequired

d-c and a-c

d-c

None

d-c and a-c

ASCS - Automatic stabilization and control system

FEW - Fly-by-wire mode Controlled by pilot

MP - Manual proportional mode \ actuation of control

RSCS - Rate stabilization control system stick

18



TABLE II.- SPACECRAFT COMMUNICATIONS AND

INSTRUMENTATION SYSTEM

Component Capability

Voice communication

UKF transceiver 2 watts

UHF transceiver 0.5 watts

HF transceiver 5 watts

Radar

CVband beacon 400-watt transponder

S-band beacon 1,000-watt transponder

Command

2 command receivers 10 channels each

Recovery

HF D/F beacon SEASAVE I watt

UHF D/F beacon SARAH 7-5 watts

UHF D/F beacon Super SARAH .... 91 watts

HF transceiver 1 watt

19



Aercxiynamic spike

Escape rocket motor

Escape tower

Spacecraftlaunch-vehicle adapter

Stub pod

Boosterengine

Spacecraft

Liquid oxygen tankpressurization line

10' dia

s

/v

\V E

-

Equipment pod

Liquid oxygen line

Vernier fairing

^Sustainer engine

Figure 4.- Space-vehicle configuration

20



: :

Roll

Pitch

P i t c h Y aw

Pitch is defi n ed as the rotation of the s p a c e c r a f ta b o u t its X - a x i s . The p i t c h a n g l e is O° w h e n the

Z-axis lies in a h o r i z o n t a l p l a n e . U s i n g th ea s t r o n a u t ' s r i g h t side as r e f e r e n c e , p o s i t i v epitch is a c h i e v e d b y counterclockwise rotationfrom the O° p l a n e . T he rate of this rotation ist h e s p a c e c r a f t p i t c h r a t e and is positive in thedirection shown.

Roll

Roll is d e f i n e d as the rotation of the spacecrafta b o u t it s Z-axis. C lockwise rotation of thes p a c e c r a f t , a s v i e w e d from b e h i n d th e a s t r o n a u t ,is called right roll and is defi n ed as positive.Whe n t h e X - a x i s o f t h e s p a c e c r a f t lies in ah o r i z o n t a l p l a n e , the roll angle is O .

Ya w is d e f i n e d as rotation of the s p a c e c r a f t a b o utits Y-axis. C l o c k w i s e r o t a t io n of the s p a c e c r a f t ,

wh en v i e w e d from a b o v e t h e a s t r o n a u t , i s calledright y a w a n d i s defi n ed a s p o s i t i v e .

Ya w angle is considered O w h e n t h e s p a c e c r a f tis in n o rm a l orbital posit ion ( bl unt end ofs p a c e c r a f t faci n g line o f f l i g h t ) . W he n th e

p o s i t i v e Z-axis of the s p a c e c r a f t is directedalong the orbital f l i g h t p a t h (s m a l l end ofs p a c e c r a f t f a c i n g line o f flight) , the yaw a n g l eis 180°.

A c c el e ro m e t er P o l a r i ty

With the spacecraft in the launch posit ion theZ-axis wil l b e p e r p e n d i c u l a r to the earth's s u r f a c ean d th e Z-axis accelerom eter w il l read +lg.

Figure 5.- Spacecraft axis system.

21



Figure 6.- The MA-6 Mercury spacecraft and launch-vehicle adapter,

22



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28





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LAUNCH-VEHICLE DESCRIPTION

The launch vehicle employed to accelerate the Mercury spacecraft intoan orbital trajectory is an Atlas Series D missile modified for the Mercury

mission. The Atlas is a 1-stage liquid-propellant launch vehicle, and

derives the greater portion of its 360,000 pounds of thrust from one sustainerand two booster engines. In addition, two small vernier engines are includedto provide precise control of launch-vehicle attitude and angular rates. Theintegral propellent tanks which contain the RP-1fuel and liquid oxygen, inaddition to providing structural rigidity through pressurization, are dividedby a pressure bulkhead. Guidance during powered flight is provided by a combi-

nation of ground-generated radio commands and spaceborne inertial referencesystems. All launch-vehicle parameters which are critical to its operationduring launch are monitored by an abort sensing and implementation system(ASK) which initiates the escape sequence in case of an impending disaster.

Airframe

The major divisions of the airframe are the tank section and thebooster-engine fairing section. The tank section is a monocoque structurefabricated from thin stainless steel and is closed at the forward end by adomed bulkhead and at the aft end by a truncated-cone bulkhead. The interme-diate bulkhead is a hemispherical structure insulated below on the fuel tankside by a layer of plastic. The insulation bulkhead retains the insulationbelow the intermediate bulkhead. The booster section structure essentiallycomprises a cylindrical shell, nacelles, fairing, and heat shield withinwhich are supported ths subassetnblies of the booster-stage propulsion,pneumatic, electrical, and hydraulic systems. The booster section isjettisoned during inflight staging.

Propulsion System

The Atlas propulsion system is designed as a fixed thrust, bipropellantengine cluster, and the system includes thrust-chamber gimbaling to providelaunch-vehicle attitude control. Each of the two booster engines developsa sea-level thrust of 15 ,000 pounds, and the sustainer engine generates avacuum thrust of 57,000 pounds. Each of two diametrically opposed engines,located just above the booster-stage separation plane, supplies up to

1,000 pounds of thrust to provide vernier control of launch-vehicle attitudesand rates. Centrifugal-type turbopumps feed fuel and liquid oxygen into thesustainer and booster engines.

Guidance System

In the early portion of powered flight, the launch vehicle is guidedby a programed autopilot; and soon after the tracking facilities associatedwith the General Electric MOD III radio guidance system acquire the vehicle, itis guided by transmitted ground commands. Tracking data are processed through

38



ri •': :." :*• • •'• •'• •" : :•• : :•• :-,••• .II I * • • ••»•••••4

- :- :•: .-• ..* : i i :..

a Burroughs computer at Cape Canaveral, and the real-time computations form

the basis of discrete error signals transmitted to the launch vehicle forsubsequent trajectory corrections. A tracking complex designed by GeneralDynamics Corporation ar i d referred to as Azusa provides a backup to theGeneral Electric-Burroughs system. Each tracking system employs the passivetechnique, in which both a rate and a pulse beacon are located in the launchvehicle for interrogation. A decoder installed in the launch vehicle receives

pulse messages and verifies and converts them into control commands whichare sent to the flight programer.

Abort Sensing and Implementation System

An automatic abort sensing system to initiate rapid spacecraft sep-aration from the launch vehicle is located within the launch vehicle todetect malfunction of the launch-vehicle systems. The ASIS monitors criticalparameters in the operation of the launch vehicle in the following areas:flight control system, tank pressurization system, electrical system, pro-pulsion system, and sustainer hydraulic pressure. In the flight controlsystem, rates about all axes are instrumented, and in the electrical system,the spacecraftlaunch-vehicle interface and 400-cycle voltage are monitored.The fuel injection manifold pressure is sensed in the propulsion system.

Aerodynamic Load Criteria

The maximum wind velocity allowable for the launch of a manned Mercuryspacecraft on the Atlas launch vehicle is 18 knots. The design criteria forthe spacecraftlaunch-vehicle combination up to booster staging include loadsimposed by the launch-vehicle acceleration and wind gusts. These designcriteria are listed below:

Dynamic pressure, q, Ib/sq ft 1,000

desr lbAngle of attack times q, oq, =--77 7,500

sq it

Normal load factor, g 0.5

Axial load factor, g 8

Gusts:

Equivalent airspeed at altitudes below Uo,000 feet,ft/sec 30

True airspeed at altitudes above 0,000 feet,ft/sec 60

Launch-Vehicle Modifications

The MA-6launch vehicle, Atlas no. 109-D, was modified for the missionas on previous Mercury-Atlas flights. It differed from the MA-5 Mercury-Atlaslaunch vehicle (93-D) in one major respect. The insulation and its retainingbulkhead between the lox and fuel tank dome was removed prior to launch when

39



It was discovered that fuel had leaked into this insulation. The originalrequirement for this insulation and retainer had been deleted earlier inthe Atlas development program as being nonessential. The following minormodifications were incorporated into the Atlas 109-D for the MA-6 mission:

1. The gyro canister was modified to include specially selected

transistors of an earlier design that displayed good thermal characteristics.This change decreases the possibility of thermal run-away in the gyro torquerand signal amplifier.

2. The launch-vehicle lox-tank pressure parameter for the abort sensing

and implementation system was changed from 21.5 ± 1-0 psi to 19-5 ^ 1-0 psito protect against an inadvertent abort resulting from lox-tank ullage-pressuretransients which occur at lift-off.



MISSION OPERATIONS

The various ground operations required to support a Mercury orbitalmission successfully may be grouped according to appropriate mission phases:prelaunch, launch, flight, and recovery. The prelaunch operations include

the preparations necessary to bring the spacecraft, launch vehicle, astronaut,and ground support personnel up to flight-ready status. The launch operationsbegin with the countdown, where all flight systems and flight-control stationsare checked for readiness, and concludes with insertion of the spacecraft intoits orbital trajectory. The orbital portion of the flight entails the flightmonitoring and data acquisition operations of personnel stationed along theMercury Worldwide Network. The recovery operations begin when a landing pointis predicted by appropriate network stations and involve the combined effortsof thousands of Department of Defense personnel stationed at the variouspre-scribed landing locations along the orbital ground track. For a brief back-ground regarding Mercury mission operations, see reference 5-

PRELAUNCH OPERATIONS

The prelaunch operations consist of the training of the astronaut fora specific flight, preparations conducted at the launch site for the spacecraftand the launch vehicle, and flight safety reviews. Although each astronauthas received training since his introduction into the Mercury program, he mustbe specially trained for the mission involved. This training primarily involvesparticipation in a series of mission simulations which present realistic oper-ational situations to be assessed and acted upon. These simulations are oftenconducted in conjunction with the detailed checkout operations completed for the

spacecraft, launch vehicle, and the Mercury Network. Program managementpersonnel attend scheduled review meetings to evaluate the status of prelaunchpreparations for the spacecraft and launch vehicle and to initiate remedialaction as necessary in order to insure the safety of the astronaut throughoutthe mission. The following paragraphs outline the operations required inpreparation for launch, beginning with the arrival of the spacecraft at thelaunch site.

Astronaut Training and Preparation

The astronaut training program for Project Mercury can be broken down into

six basic categories whichare

essentially dependenton the

training devicesused. These categories are academics, static training, environmental familiar-

ization, dynamic training, egress and survival training, and specific missiontraining. For a more complete description > a n d history of Mercury astronauttraining, see references 6 and 7

and paper 10 of reference k.

Academics.- In addition to a series of lectures on basic astronautics earlyin the program, detailed systems briefings were given by contractor engineersconcerned with the design of the various systems. Also, NASA engineers con-cerned with the monitoring of these design tasks presented detailed briefingson their specific spacecraft system and current changes involved. As a

i n



supplement for classroom lectures, a number of field trips were made to con-tractor plants where flight hardware was developed. Each of the astronautswas assigned a specialty area associated with a phase of Mercury; and in thecapacity of consultant for their particular area, the astronauts participatedin and contributed to design reviews from an operations standpoint.

Static training devices.- Soon after the beginning of the astronaut

training program, a series of training devices were designed and made avail-able. Initially, rudimentary attitude-control trainers were employed. Thesesimulators were soon replaced by the more sophisticated procedures trainers.At present, one of the two procedures trainers is located at Manned SpacecraftCenter, and the other is installed at the Mercury Control Center, Cape Canaveral.These simulators provide valuable training in all mission phases, and many ofthe possible failure modes may be simulated and the resulting performance ofthe astronaut may be evaluated. In addition to providing valuable trainingfor the flight astronauts during critical mission phases, these static trainingdevices have allowed accurate assessments of system changes affecting pilot pro-cedures which have been incorporated into the trainers.

Environmental familiarization.- A basic phase of preparing the astro-nauts for flight is their familiarization with the extreme environmental con-ditions which might occur during a Mercury mission. Two examples of specificenvironmental training are the thermal chamber and the carbon dioxide room.The thermal chamber provided exposure to maximum expected heat stresses duringa simulated reentry. The carbon dioxide room merely offered some familiaritywith the effects of above normal concentrations of this gas in order to makethis condition more easily recognizable. In addition, the pilots have beenconditioned to the normal spacecraft environment through their participationin altitude chamber tests as a part of the prelaunch preparation of thespacecraft.

Dynamic training devices.- The training units which can be classifiedas dynamic in nature include the zero-gravity trainers, the centrifuges, theair-lubricated free-attitude (ALFA) trainer, the multiaxis spin-test inertiafacility (MASTIF), and various high-performance jet aircraft. The zero-gravitytrainers are primarily high-speed transport aircraft, which can provide up to30 seconds of weightlessness by flying a parabolic arc. In addition, valuabletraining in this regard was also achieved through hydrostatic test facilitiesand scuba diving. The Aeromedical Acceleration Laboratory (AMAL) located atJohnsville, Pennsylvania, includes a centrifuge which permits simulations in

mission phases, such as launch, reentry, and abort, during which the acceler-

ation levels are particularly high. A mock-up of the spacecraft interior isinstalled in the gondola of the Johnsville centrifuge, and after depressur-ization to nominal cabin pressure for a Mercury flight, the simulated missionsin this trainer are exceptionally realistic. A major dynamic trainer is theALFA trainer. A completely enclosed couch is mounted on an air bearing; andwith the Mercury hand controller which actuates compressed-air jets, thetrainer attitudes can be controlled about all three axes. Simulation of theretrorockets has been provided by additional compressed-air jets, and thedifficult manual retrofire maneuver can be effectively practiced. The MASTIFtrainer is similar in function to the ALFA trainer, but it is mounted so that



• •*

rapid rates about all three axes can be generated for a glmbaled contour seatand hand-controlled configuration. Practice in recovering from violent space-craft maneuvers is effectively conducted. And finally, each astronaut hasmaintained proficiency in operating high-performance aircraft, primarily asupersonic jet airplane assigned to the program.

Egress and survival training.- Since the primary recovery area forMercury is in the water, many hours have been spent in practicing egressfrom the floating spacecraft. The primary mode of egress is through thesmall pressure bulkhead and the recovery section. In most cases, maximumutilization has been made of actual recovery conditions and vehicles, suchas helicopters. Survival training was provided in various degrees of hard-

ship, including -~ day in a liferaft and 3 days in the desert.

Specific mission preparation.- Although the training activities for theastronauts have been reasonably concentrated during the 3 years of initialpreparation, specific assignments immediately prior to a flight for the chosen

pilot are especially demanding. For the MA.-6 astronaut, these assignmentsincluded complete familiarization with spacecraft 13, physiological examina-tions (discussed in the Aeromedical Analysis section), flight simulations(discussed in the Astronaut Flight Activities section) using the actual flightplan, physical conditioning, and familiarization with special equipment to beincluded and used during the flight. In addition, more specific tasks in thetraining devices described previously were completed. The flight and backupastronauts also actively participate in the many spacecraft and launch-vehiclesystems reviews in the days immediately prior to launch. The pace of activitywith regard to precise astronaut training steadily increased throughout theprogram as new knowledge was gained and operational demands became more severe,and the MA.-6 flight was the culmination of these efforts.

Spacecraft Prelaunch Preparation

The detailed checkout of the spacecraft is conducted primarily at CapeCanaveral during the prelaunch preparation period. The activities duringthis period are additionally described in reference 8 and paper 5 of referenceThe detailed examination involves functional testing of the spacecraft systems,observing in detail the performance of the individual systems. These tests,which duplicate as nearly as possible the different flight environments andmodes, are repeated as necessary. During these tests, all discrepancies, nomatter how trivial, are scrutinized for their significance. Design changes

indicated by these tests and the previous Mercury flights are incorporated asrapidly as possible so that the optimum spacecraft configuration may be flown.Astronaut participation in all system checkouts and design reviews at CapeCanaveral is very important to the mission. This participation provides thepilot with intimate familiarization with the flight spacecraft and its systems.

Time utilization.- The checkout phase for spacecraft 13 was conducted inHangar S, assigned to the NASA, at Cape Canaveral, Florida, and lasted approxi-mately 133 days as shown in figure 19- This checkout is followed by launchcomplex operations in which the launch vehicle and spacecraft are mated and



• • •• • ••• • •• • •••

tested together. This testing is conducted to assure that the launch vehicleand spacecraft are mechanically, electrically, and radio-frequency compatibleand is followed by final assembly and launch preparations.

The time spent in the hangar can be initially broken into two parts:

(1) Work performed on the spacecraft, such as spacecraft assembly andservicing and incorporation of design changes. For MA-6, this worktook approximately 63 of the 133 days spent in the hangar.

(2) Systems testing, troubleshooting, and replacement of components asa result of this testing. Seventy days were spent in this phase ofthe operation.

The launch-complex operations required 33 days to complete.

Design changes. - The MA-6 spacecraft had a total of 255 minor designchanges incorporated at Cape Canaveral, a portion of which resulted from the

experience gained during the orbital mission of spacecraft 9 (MA-5). Thesechanges involved the following systems: the reaction control system, theelectrical system, the ECS, the ASCS, and items of a miscellaneous nature.The major changes are presented in the Spacecraft History sections.

Systems tests.- Checkout operations at Hangar S consist of individualsystems tests followed by a simulated flight test with all systems operatingin a manner approaching flight conditions. The following system tests wereperformed on spacecraft 13:

(1) Electrical power

(2) Instrumentation(3) Sequential(U) Environmental control(5) Communications(6) Reaction control(7) Communications radiation tower test(8) Automatic stabilization and control(9) Altitude chamber tests

The electrical power system test was the initial test performed on theMA-6 spacecraft. This test determines if power can be safely applied to thecontrol and power distribution system of the spacecraft. The test also checks

automatic and manual a-c inverter switching. Power surges on the d-c bus wereexperienced when the 150 v-amp main inverter was switched on the line. Becausean excessive number of power surges occurred, the inverter was replaced.

The instrumentation system was thoroughly tested and calibrated on thebench prior to its installation in the spacecraft. The system was tested againafter it had been installed in the spacecraft. The primary purpose of thelatter test is fourfold: (l) to determine the error, if any, between thehardline signal normally used to transmit data to the blockhouse or other testequipment in the hangar and the signal radiated through the telemetry trans-mitter* (2) to determine possible interference within the instrumentation



system and "betveen it and other systems operating in the spacecraft; (3) tomake a single-point check of the complete calibration made on the "bench; and(4) to insure that the system is satisfactory to support other spacecrafttests. No major system discrepancies were uncovered during this testing.

The sequential system test provides for the checkout of the automatic andmanual sequential system. The sequential system testing may "bedivided intothe following four major parts: launch, orbit, escape, and reentry. Con-trolled inputs are fed to the sequential system and the system outputs aremonitored. The maximum-altitude sensor actuates the jettisoning of the escapetower. Although called a maximum-altitude sensor, it is actually a variabletimer whose time delay depends on the existence of an abort signal and the timefrom lift-off at which the abort signal occurs. During the abort phase in thesimulated flight of spacecraft 13, it was determined that this sensor actuatedprematurely. The sensor was replaced and the abort runs were repeated suc-cessfully.

The environmental control system (ECS) checkout, which is conducted priorto the altitude-chamber test, determines the functional operation of theseparate components of the ECS system. The oxygen bottles are serviced tooperating pressure at this time and maintained at this level for the altitude-chamber test.

As the result of this testing, the following major ECS discrepancies wereuncovered:

(1) Excessive leakage of the high-pressure oxygen shutoff valve wasnoted.

(2) The high-pressure oxygen regulator showed an external leak. Thevalve body was found to be defective.

(3) An aluminum check valve for the Freon coolant supply had stuck inthe open position and all valves of this type were replaced withstainless-steel valves.

The communications systems tests are conducted in two parts, the benchtests and the radiation tower tests. The primary purpose of the bench testsis to determine the electrical characteristics of the individual componentsthat make up the onboard communications system. For the radiation tower tests,

the flight configuration is simulated as closely as possible. These testsevaluate the transmission quality of the HF bicone antenna; other communicationssystem components are tested at the same time. No significant discrepancieswere revealed during the systems tests for the MA-6 spacecraft.

The reaction control system (RCS) checkout procedure determines thecondition and operation of the RCS. These tests are conducted in a specialtest cell, and several gas checks are employed to determine the overall gasintegrity of the RCS, as well as functional tests of the entire system. Thefinal test of this series involves the static firing of the pressurized RCS.The system performed well during the testing of spacecraft 13.



' ' •

The automatic stabilization and control system (ASCS) checkout is dividedinto two parts, a static and a dynamic test. The dynamic test is conductedwith the spacecraft in a fixture which is rotated at a constant rate in pitchand roll. For MA-6, the static test was completed without incident, hut thedynamic portion of the test was interrupted by a failure of the yaw-repeater

loop. A shorted capacitor was discovered in the repeater motor circuitry. Anew autopilot passed both the static and dynamic tests with no discrepancies.

The altitude-chamber tests are used to determine the peculiar operatingcharacteristics of the overall environmental control system (ECS) for theflight. The astronaut is suited and connected into the ECS for the first timeduring these tests. The chamber is pumped down to a simulated altitude ofabout 125,000 feet, and a simulated flight is initiated. MA-6 teat data indi-cated that the 150 v-amp inverter overheated during the first three runs, andthe water-flow orifice to the cooling plates was increased in size. Furthertesting revealed that the sealant used in the construction of the coolingplates had penetrated into the flow passages and clogged them. The passages

were cleaned and the original orifices were installed. Flow tests on the finalinstallation showed that the system was clean.

Simulated flights.- Following the systems tests, a series of simulatedmissions is begun. The simulated flight test for spacecraft 13 began inHangar S on November 25 and was completed on December 12, 1961. When thelaunch was rescheduled for January 19 2, another hangar simulated flight wasinitiated on December 19 and was successfully completed on December 21, 1961.At this time, the spacecraft was considered functionally ready for launch padoperations. For greater detail regarding the technical objectives of thesimulated flights and their results, see pages 6l and 62 of reference 4.

Launch pad operations (prior to countdown).- The launch pad operationsprior to countdown are scheduled for completion in-13 days; however, becauseof delays caused by changes in the MA-6 spacecraft and launch vehicle aftermating, changes prompted by trouble encountered during tests, and adverseweather conditions, the actual time the spacecraft spent on the launch padwas somewhat longer than this, as shown in figure 20. The launch pad oper-ations prior to countdown are listed below in the order in which they arenormally performed:

(l) Launch complex checkout2) Interface inspection

3) Mechanical matingh) Spacecraft systems test5) Electrical interface and aborts6) Flight acceptance composite tests(7) Flight configuration sequence and aborts(8) Launch simulation, including RCS static firings(9) Simulated flight

(10) lyrotechnic check(11) Spacecraft servicing(12) Precount



• ••••

Each of these operations, together with the troubles encountered, subse-quent delays, and repairs made, is discussed in greater detail in reference 4

(pages 62 to 66).

It is apparent that the detailed preflight testing is a time-consumingprogram; however, this testing has revealed a number of system malfunctionsor weaknesses which might have prevented the completion of a successful mis-sion, not only for MA-6, but for previous flight tests as well. Although thispolicy of detailed testing and checkout has led to lengthy preparation periods,inflight failures would cause even longer delays in the overall program.

Spacecraft History

Spacecraft no. 13 was delivered to Hangar S at Cape Canaveral, Florida,on August 27, 1961. Upon arrival, the spacecraft entered a brief period offinal installation, during which the individual systems checks were completedsatisfactorily prior to preparing the spacecraft for the altitude chamber tests.

Reaction control system.- The following items represent major changes tothe reaction control system during testing at Cape Canaveral prior to launch:

(1) Roll thrust-chamber heat sinks were added.

(2) Existing flare seals were removed from the automatic system inletand outlet connections of the thrust-chamber solenoid valves toreduce the possibility of leakage in the area of the thruster. Sub-sequently, all applicable seals were replaced with an improved seal.

(3) Nine solenoid valves were replaced with aluminum valves because ofthe inferior quality of the plastic seals at high-temperature levels.

(4) The clockwise automatic roll-thruster assembly was replaced.

Environmental control system.- During the testing performed at CapeCanaveral, the following changes in the environmental control system were made:

(1) Water-type heat exchangers of an advanced design were installedunder the 150- and 250-v-amp main inverters.

(2) Screens with 0.06-inch diameter holes were installed in the cabin

fan inlet ducts.

(3) The aluminum check valves in the cooling system for the inverterswere replaced with stainless steel valves.

The high-pressure 0 reducer was replaced because of a leak prior

to-the altitudechamber runs.

(5) The primary-0 -system shutoff-valve stem was removed, reworked, and

reinstalled because of leaking 0-rings.

^^^M__CONFIDENTIAL



••t ••

(6) A manual "blood-pressure system vas installed.

(7) During the first run of an entire simulated mission, the no. 1 suitfan was "below specification performance and was replaced.

(8) An indicator was installed on the instrument panel to provide a

readout of suit-circuit steam-vent exhaust temperature.

(9) The CO absorber vas replaced after the spacecraft was moved to the

launching complex.

(lO) The suit-circuit demand regulator was replaced.

Communications.- No major problems were experienced with the communica-tion equipment after the spacecraft arrived at Hangar S. Readouts from therange on the C-band radar beacon showed jitter and varying signal strength.These problems disappeared, as expected, when the service tower was removed.

Electrical system.- The following changes were made to the electricalsystem at Cape Canaveral:

(1) Flight batteries were activated on November 8, 1961. These batteries,which were near their normal activated lifetime, were replaced onFebruary 10, 19&2, when one cell lost voltage and another had a shortcircuit.

(2) The mercury-cell auxiliary battery units used for the integratingaccelerometer and the maximum-attitude sensor were replaced withnew units on January 13, 19 2.

(3) Inverter switching circuitry was modified to provide a nonfusedcircuit to the standby inverter in the manual mode.

(k) The original 150 v-amp inverter malfunctioned during a hangar testand was replaced.

(5) The suit-fan toggle switch was replaced.

(6) The removable half of all fuse-block holders was reinforced.

(7) The telelight panel was removed, reworked, and reinstalled.

(8) The retrorocket relay panel was replaced.

(9) Because of an internal malfunction, the thrust-cutoff sensor wasreplaced.

(10) During a hangar check, the maximum altitude sensor actuated pre-maturely and was replaced.

(11) The satellite clock was replaced on two separate occasions becauseof malfunctions.



.••

(12) Indicator lights shoving which inverter was operative were added tothe instrument panel.

Launch-Vehicle Prelaunch Preparation

In order to support Mercury-Atlas missions, "both the launch complex and

the launch vehicle required modification. Figure 21 shows the major modifica-tions made to the launch complex. In the service tower, a room was built toenclose the spacecraft. This room, commonly called the "white room," is lo-cated near the top of the service tower. The environment in this white roomis controlled to minimize the effects of humidity, dust, and so forth, on thespacecraft components. An emergency egress tower is shown in figure 22. Thefigure depicts the egress platform in the extended position, such that theend of the platform is adjacent to the hatch of the spacecraft. When retracted,the platform arm is rotated in the vertical plane so that the launch vehiclewill not strike the platform when it is launched. Special rescue and fire-fighting vehicles are stationed just outside the complex to transport the

egress crew to the tower and/or to meet the astronaut at the tower and trans-port him away from the complex. A special firefighting system including fournozzles is also installed and is remotely controlled from the blockhouse insuch a manner that water or fire-smothering foam could be directed to any areainside the complex. A radio command link is provided in the blockhouse. Thiscapability provides a ground-command means of firing the spacecraft escaperockets and aborting the spacecraft prior to launch and is the primary meansfor providing an abort during the first 10 seconds of flight.

Upon arrival at Cape Canaveral, the launch vehicle is inspected and pre-pared for erection in approximately 48 hours. The launch vehicle is trans-ported to the complex on a dolly-type vehicle, backed into the launcher, alined,

and attached to the launcher. A hoist cable is then attached to the front endof the dolly (top end of the launch vehicle) and the dolly and launch vehicleare hoisted to the vertical position. The launcher is then rotated back toits original horizontal position.

For MA-6, it was learned after erection that the launcher mechanism couldnot be adjusted sufficiently to aline the launch vehicle properly. Therefore,the launch vehicle was taken down, the launcher mechanism was replaced, andthe launch vehicle was reerected.

All systems on the complex and the launch vehicle are then tested indi-

vidually. Complete tanking tests are conducted in which the fuel and liquidoxygen tanks are loaded and pressurized to flight pressure. This test isperformed to determine if any leaks are in the systems and also to check outthe controls related to each system. During this test on the MA-6 launchvehicle no major leaks were evident; however, some minor leaks were discoveredand subsequently corrected.

The autopilot system is also completely tested; however, before autopilottests are conducted on the launch complex, the gyro packages are calibratedin a laboratory at Cape Canaveral with special testing equipment. Thesepackages are also electrically mated to the abort-sensing control package

49



:"• . j • • ' • ' :. : :' i - 5 : 5• • •• ? * ••• ••

•••••CONFIDEHTIAL

which is part of the ASIS previously discussed. During these laboratory andsystems tests, various anomalies vere uncovered in the gyro package and theASIB control package. These packages vere replaced and systems tests werecompleted satisfactorily on the launch vehicle.

All launch-vehicle systems are then tested simultaneously in a critical

test commonly knovn as the flight acceptance composite test (FACT). Thistest is conducted to determine that all launch-vehicle systems are compatible.The FACT must be successfully accomplished before the spacecraft is elec-trically mated to the launch vehicle. After electrical mating, the launchvehicle and spacecraft participate jointly in all tests. For a more completedescription of the launch complex and the preparation for the launch vehicle,see paper k- of reference k.

The first attempt to launch MA-6 was on January 27, 1962. The launchvehicle was loaded with fuel on January 2^. However, the mission was canceledbecause of excessive cloud cover in the launch area and was rescheduled for

February 1. The fuel tank was drained, and on January JO, it was again loaded;however, normal inspection procedures disclosed that the insulation-retainingbulkhead in the fuel tanks was leaking. The leak was in the lower retainerand had allowed fuel to soak into the insulation and become trapped. Aftera careful study of the possible resultant effects, it was decided that suffi-cient flight experience had been obtained to justify removing the retainerand the insulation. After the retainer and insulation were removed and allsystems were reconnected, a complete test program was rerun on every systemdisturbed by the modification. The simulated flight test was rerun onFebruary l6, 1962, and the launch vehicle was again loaded with fuel in prep-aration for launch on February 20.

Flight Safety Reviews

Two series of meetings were held by the MA-6 Flight Safety Review Boardbecause of the launch postponements.

First series of reviews.- The meetings were conducted in anticipation oflaunch on January 24 and again on January 27, 192. The launch was rescheduledfrom January 2^ to January 27 when an oxygen leak in the spacecraft environ-mental control system was discovered on January 20. The countdown on January 27was conducted until 20 minutes before launch, at which time the weather causedpostponement of the launch attempt.

The first spacecraft review meeting was held on January 18. The space-craft history at AMR and the present status of all the spacecraft systems werereviewed, after which the spacecraft was approved as ready for flight. Asecond review meeting, scheduled after the oxygen leak had been repaired, washeld on January 2^. During this meeting, the status of the spacecraft systemswas discussed, and all systems, including the ECS, were again approved asready for flight.

The first Booster Review meeting was held on January 19. All launch-vehicle and supporting systems were approved as ready for flight. The

50



• ••• ••• •

Mission Review meeting vas held at 1:00 p.m. on January 25, and all elementsfor the flight were found to be in readiness.

When the Flight Safety Board convened at 9:30 a.m. on January 26, itwas reported that the launch vehicle was not in a flight status because ofbroken wires and damaged pins in a separation plug. This plug, which is

disconnected at booster staging, carries booster engine autopilot commands,rough combustion cutoff circuitry, and engine instrumentation signals. Theplug was repaired, and,when the Flight Safety Board met again at 1:00 p.m.on January 26, the launch vehicle was ready to be committed to flight. Weatherproblems caused the launch attempt of January 27 to be postponed.

Second series of reviews.- On January 30 fuel was discovered in the in-sulation between the structural bulkhead and the insulation bulkhead separatingthe launch-vehicle fuel and oxidizer tanks. Fuel had leaked into this areaaround a flange bolt. The decision was made to remove the insulation andinsulation bulkhead, and this work period caused the launch to be rescheduledfor February 13. Adverse weather forced three more postponements to

February lk, 15, and finally 20. No separate spacecraft or launch-vehiclereview meetings were held during this period. The second Mission Reviewmeeting was held on February 12. Satisfactory removal of the insulation andretaining bulkhead had been made, and all systems were found to be ready forflight. Launch-vehicle status meetings were held on February 13 and 19, andthe Flight Safety Board recommended that the mission proceed since all systemswere in readiness.

CONFIDENnAL 51



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LAUNCH OPERATIONS

Launch Procedure

The spacecraft launch operations were planned incorporating a 6lO-minute

split countdown with a 17 -hour built-in hold at T-390 minutes for spacecraft

peroxide servicing. In order to provide additional assurance that the time

of lift-off would not be later than the preestablished m a x i m u m time of10:00 a.m. on February 20, 1962, a 90-minute hold was scheduled at T-120 minutes.

The second half of the split count was picked up at 11:30 p .m., e.s.t.,

on February 19, 1962. Launch was at 9:^7 a.m., e.s.t., on February 20, 1962,

after 2 hours and 17 minutes of unplanned holds. The following is a sequence

of important events, including holds, which occurred in the countdown:

T-J90 min Count was resumed.

T-120 min Built-in 90-minute hold. Because of a sudden drop inthe automatic gain control of the launch-vehicle rate

beacon, the first backup beacon was substituted forthe original during this 90-minute scheduled hold

period. The hold was extended an additional ^5 minutesto complete installation and revalidate the beacon.A n additional 10-minute hold was required to replace

a broken microphone bracket in the astronaut's helmet.

T-120 min The count was resumed.

T-117 min Completed second launch-vehicle guidance loop test.

T-87 min Hatch installation began.

T-60 min A ^0-minute hold was required to replace a brokenhatch bolt. A third launch-vehicle guidance loop

test was performed during this hold.

T-^5 min A 15-minute hold was required to add approximately10 gallons of fuel to the launch vehicle.

T-22 min A 25-minute hold resulted from a malfunction of the

main lox fill p u m p outlet valve. The final 20 percentof lox tanking was accomplished by using a smaller

p u m p via a 6-inch line, resulting in a slower operation.

T-6 min JO sec A 2-minute hold was required by the Mercury flightdirector to investigate a loss of power to the Bermuda

computer.

Weather Conditions

The weather conditions in the launch area at lift-off were as follows:p

Cloud cover ........................ ^n" alto-cumulus

Visibility, miles ................... .- ...... 10

Surface winds, knots ............... North at l8 (gusts to 25)

56



• • I• •

A plot of the launch area vlnd direction and speed is shown in figure 23for altitudes up to 60,000 feet.

Landing area weather conditions at lift-off were as follows:

Cloud cover

Visibility, miles ........................... 10Surface winds, knots ..................... ESE at IkWaves, feet .............................. 2

Photographic Coverage

Atlantic Missile Range (AMR) optical coverage, including the quantity ofinstrumentation committed and data obtained during the launch phase, is givenin table III. AMR optical tracking from lift-off or first acquisition to thelimit of visibility is shown in figure 2^.

Metric film were reduced and the results were tabulated by AMR, but thesedata were not required for evaluation by MSC since the poweredflight phasewas normal .

Engineering sequential coverage at AMR station 1 during the launch phasewas satisfactory. Thirteen 16 mm and 35 nun films from three fixed and tentracking cameras were reviewed. The quality of fixed camera coverage wasexcellent and indicated normal umbilical ejection, periscope retraction, launch-vehicle ignition, and lift-off. The quality of tracking camera coverage wasgood during all phases, with the exception of the early portion of poweredflight because of ground haze conditions at lift-off. All tracking cameras

indicated normal launch-vehicle staging and escape-tower separation. Docu-mentary coverage used for engineering evaluation of the mission was satisfac-tory, and film quality was average. Seven motion picture films and numerousstill photographs were available for review. Two motion picture films pre-sented a portion of the prelaunch activities, including astronaut preparationat Hangar S, insertion of the astronaut into the spacecraft, closing of thehatch and securing for launch, and portions of the operational activity atthe Mercury Control Center during the mission. Coverage and quality of thesetwo films were good. Ibur motion picture films presented portions of therecovery operation, including aerial and shipboard coverage of spacecraft re-trieval from the water, removal of the hatch, astronaut egress, transfer fromthe recovery ship, and the physical examination aboard the carrier. One film

included views of the astronaut at Grand Turk Island and the spacecraft beingloaded aboard the aircraft for removal to Cape Canaveral. Photographic coveragein single-frame exposures included views similar to those documented by motionpictures, with the exeption of hatch removal onboard the recovery ship.Numerous engineering still photographs were available showing close-up viewsof the spacecraft after recovery and during postflight inspection at Hangar S.

57



• • • •• •• • •• • • •

I ••• ••

TABLE III. - AMR OPTICAL COVERAGE OF LAUNCH AND REENTRY PHASES

Film type

Metric

Engineering sequential,station 1



Documentary

No. of items

committed15

1 * 7

1

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88

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obtained15

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0

88

Lost

items0

ai

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al

0

Reason for loss

Not applicable

No reentry data

No reentry data

No reentry data

Not applicable

rlanned for reentry coverage.

58



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It must be strongly emphasized that without a pilot in the spacecraft tomake decisions and take corrective action, the malfunctions which occurredwould have made the successful completion of the flight extremely difficult,if not impossible.

RECOVERY OPERATIONS

Recovery Plans

Figure 25 shows the Atlantic Ocean recovery areas where ships and air-craft were positioned at the time of launch. Areas 1 through 6 were availablein the event that it became necessary to abort the mission during poweredflight. Recovery forces were distributed so as to provide for recovery withina maximum of 3 hours after landing in areas 1 and U, and a maximum of 6 hoursin areas 2, 3, 5> and 6. Areas 7, 8, and 9 were available for landing at theend of orbital passes 1, 2, and 3> respectively, and recovery forces weredistributed to provide for recovery within a maximum of 3 hours. A total of2k ships and 15 aircraft were on station in these Atlantic recovery areas atlaunch time. In addition, helicopters, amphibious surface vehicles, and smallboats were positioned for recovery support near the launch site. Figure 26shows the positions of contingency recovery aircraft that were on alert atvarious staging bases in the event that a landing occurred any place alongthe orbital ground track. These aircraft were equipped to locate the space-craft and to provide local emergency assistance if required. See paper 7 ofreference h for a more detailed description of the MA-6 recovery plans.

Recovery Procedure

All recovery forces were on station at the planned launch time. Weatherconditions were favorable for location and retrieval in all primary recoveryareas and in the contingency areas. Recovery communications were good through-out the entire operation, and the recovery forces were informed of missionstatus during the launch, orbital, and reentry phases. During the third orbitalpass, recovery units in area 9, shown in figure 27, were alerted to expect alanding in their area, and at 0 : 2:00 (13 minutes prior to landing) theseunits were informed that the landing was calculated to occur at 21°29' Northlatitude and 68° 8' West longitude. This information was transmitted to therecovery forces as CAL REP 1 (calculated landing position report), as shownin figure 27- Continued radar tracking made little change in this prediction,

and at OU:U6:00 (9 minutes prior to landing), the recovery forces were directedto orient their search about this position. This information was transmittedas DATUM REP 1 (datum report) indicating that this was the best landingposition information available at that time. In the meantime, lookouts on thedestroyer U.S.S. Noa, which had been stationed in the DDl6 recovery position(see fig. 27) heard a noise like an explosion, and approximately 20 secondslater, the main parachute and spacecraft were sighted at an estimated slant

range of 5 miles and elevation angle of 35°• The noise was described ashaving been similar to that produced by the shock wave of an aircraft travel-ing at supersonic speeds.

62



" ••" :."' ::." :• • ..- : ; :

!• •

The U. S.S. Noa established communication with the astronaut, and at05:07:CO (12 minutes after landing), it was alongside the spacecraft. Theastronaut remained in the spacecraft during the retrieval operation. Aphotograph of the spacecraft prior to retrieval is shown in figure 28. A"shepherd's crook" was used to attach a lifting line to the spacecraft, whichwas hoisted clear of the water at 05:12:00 and secured on the recovery ship

at 05:15:00 (20 minutes after landing).

The astronaut first decided to egress from the spacecraft through theparachute compartment, whereupon he initiated the standard procedure of re-moving the right-hand instrument panel. However, considering the time andeffort normally associated with this egress route and the fact that he was un-comfortably hot,he elected to egress more rapidly through the side hatch.He actuated the hatch explosive mechanism from inside the spacecraft and wasclear of the spacecraft at 05:3 :00 (39 minutes after landing).

The following retrieval information was reported by the recovery ship:

Position of retrieval:North latitude 21°25.6'

West longitude 68°36.5'Wind velocity, knots 18 (119° from true north)Wave height, feet 2Water temperature, °F 81Air temperature, °F 75

Recovery Aids

All spacecraft recovery aids functioned normally. Search aircraft re-

ported making contact with both SARAH recovery beacons and with the UHFtransceiver. These aircraft were proceeding towards the calculated landingposition at this time and were well within the available range of these sys-tems. The dye marker and flashing light were reported to be functioningnormally. A fix,based upon the SOFAR bomb signal, was available at the re-covery center about 1 hour after landing. This fix was approximately k milesfrom the spacecraft retrieval position, as shown in figure 27- The SEASAVEbeacon fixes, as reported by the HF/DF networks of the Navy and the FederalCommunications Commission (FCC), are also shown in figure 27. The SEASAVEfirst fix was made available to the recovery center at 05:20:00 and a laterfix was available at 05:27:00. Both fixes were about 25 nautical miles from

the retrieval point.



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MISSION PERFORMANCE

The technical results of the MA-6mission are presented and accom-

panied by an analysis of the flight data. Performance analyses are groupedinto the following seven major areas: spacecraft, aeromedical analysis,astronaut flight activities, astronaut's flight report, launch vehicle,trajectory and mission events, and the Mercury Worldwide Network. Thespacecraft performance section treats each major system within the Mercuryvehicle and presents the flight results and postflight analyses. In thesections regarding the MA-6astronaut and his performance during the mission,a comprehensive medical investigation, a detailed analysis of his flightactivities, and a personal narrative account of the orbital flight experi-ence are documented. The section which presents launch-vehicle performanceis a very brief synopsis of Atlas systems' operation. The trajectory andmission events section consists largely of a graphical presentation of

major trajectory parameters and mission event times. The Mercury Network isanalyzed in the areas of tracking, communications, and computation, and theresults of the Network performance are compared with those of previousorbital flights.

68



SPACECRAFT PERFORMANCE

The spacecraft as an entity performed adequately. Some malfunctionswere experienced, and analyses of these are presented in the followingparagraphs. Also discussed, from an overall mission viewpoint, are thespacecraft systems' general performance. Flight data and measurements aregenerally not shown, other than those which clarify an analysis or presentmeasurements of particular interest. Complete time histories of spacecraftdata, without analysis, are presented in part II of this report.

Spacecraft Control System

With the single exception of " b o t h 1-pound yaw thrusters failing, thespacecraft control system functioned normally throughout the flight. Dis-crepancies reported by the astronaut between the attitude indicators andthe visual reference resulting from procedural problems are discussed, anda detailed analysis of the thruster malfunction is presented in the Reaction

Control System section.

System description.- The spacecraft is capable of the followingcontrol modes:

1. Automatic stabilization control system (ASCS), with orbit mode,orientation mode, and auxiliary damping mode.

2. Fly-by-wire (FB¥) manual,

3. Manual proportional (MP).

4. Rate stabilization control system (RSCS).

Modes 1 and 2 employ the automatic reaction control system (RCS)while modes 3 and k use the manual RCS. Each reaction control system isindependent of the other. Combinations of modes 1 and 3, 2 and 3, or 2and k may be used simultaneously. The amplifier-calibrator (Amp Cal) in-corporated a new single-pulse, orbit-mode logic, thus eliminating themultiple pulses that have been experienced previously in the orbit modebecause of dirty repeater sectors. In addition, the horizon scanner refer-ence levels were set at approximately 25 percent to lessen cold cloud

effects, and scanner slaving was programed for 8.5 minutes in each 30-minute

period.

Flight description and analysis. -

Powered flight and turnaround: The control system operation was normalduring the launch phase, but the 5 seconds of separation rate damping was

delayed 2. 5 seconds by the sequence circuitry associated with the 0. 20gswitch; thus a fairly large initial roll error was produced at the start ofturnaround. The source of this error was the preclusion of rate damping asa result of reopening the 0. 20g relay when posigrade thrust acceleration wassensed. Turnaround was managed adequately, although the time required

69



••I •

(38 seconds) to settle into orbit mode was longer than normal because of theinitial roll error. Spacecraft attitudes and rates near insertion arelisted in the following table:

Event

Separation

Separationplus 2.5 sec(start ofdamping)

Separationplus 7 sec

(end ofdamping andstart ofturnaround)

Roll

Attitude,deg

0

-12

-22

Rate,

deg /sec

-1

-6

0

Pitch

Attitude,deg

-8

2

4

Rate,

deg /sec

0.4

5 - 5

0

Yaw

Attitude,deg

180

170

160

Rate,

deg /sec

-0.3

-3.0

0

Orbital phase: Except for the thruster malfunction, the controlsystem during the orbital phase functioned essentially as designed. Con-trol system exercises and usage modes are discussed in a later section. Withthe initiation of the first yaw maneuver, attitude indicators began to dis-

agree with true spacecraft attitudes. Such disagreements are inherent inthe system and occur whenever the yaw or roll attitudes deviate from 0° foran extended period of time, which can be demonstrated both mathematicallyand with operational hardware. Values differ, depending upon the presenceof either scanner slaving or fixed-pitch orbital precession, but the effectsare similar. Each disagreement has been examined and can be explained foreach system design.

The following hypothetical maneuver can best illustrate this point:

Assume that the spacecraft and gyros are properly erected in the normal

orbit attitude. The spacecraft is then yawed 90° and maintained in this yawheading for r- of an orbital period (11.25 minutes or 45° orbital travel),

during which time the astronaut maintains local vertical using the window

reference. During this •& pass, the fixed-pitch precession signal (about

k°/min) will drive the vertical gyro spin axis in a direction 90° from theorbital plane, and at the end of this period the pitch attitude indicatorwill disagree with the spacecraft true attitude by 45°. At the same time,the astronaut will have rolled the spacecraft ^5° to maintain his verticalposition with respect to the local horizon. If the spacecraft is then

70



•• ... . J99 9

restored to the normal orbit attitude, the gyro spin axis will maintain itsposition in space. This results in permanent attitude errors in both axes,unless corrected by slaving.

The 180° yaw maneuver, initiated at QJilk-.QQ is offered as the bestexample of the preceding hypothetical maneuver (see fig. 29). Indicated

attitude errors, produced by spacecraft maneuvering, can be avoided byleaving the gyros caged during such maneuvers. New gyro references canbest be restored by continuous scanner slaving, combined with uncaging thegyros at approximately 0° pitch and then permitting the spacecraft to remain

at orbit attitude (-3 ° pitch) for approximately 5 minutes.

Three gyro cagings were executed by the astronaut during the orbitalphase of the mission. In each instance the spacecraft was alined by pitchingdown to -15° to bring the horizon into view through the window. During thefirst caging the spacecraft roll attitude was -19° •

Horizon scanner operation was as expected, with short roll scanner"ignore" periods occurring just prior to sunset because of saturation of thesensors. All ignore signals produced by the pitch scanner can be correlated

with maneuvers. Eight programed cycles of horizon scanner slaving werecom-pleted which promptly removed all maneuvers and caging errors.

Retrofire: Retrofire began at 0 :33:08 and was completed by 0 :33:33.

The astronaut provided backup to the ASCS by using manual proportional control

and maintained slightly better attitudes than were experienced on theMA-5mission. Pitch and roll attitudes were held to within 1°, and the yaw devia-tion did not exceed 2°.

Reentry: The pitch -up maneuver to reentry attitude was initiated by theastronaut at OU:39:39> an<

i the 0.05g relay was actuated manually at OU:U3:31-The planned roll rate for reentry was manually initiated at 04: 4:41, and thisrate reached a value of -ll°/sec within 2 seconds. Spacecraft oscillations,with a period of 1-5 seconds, were evident about the pitch and yaw axes byOk:k6:J>0. An attempt was made by the astronaut to damp these oscillations.By 0 : 7:17, the rate of oscillation in pitch had increased to greater than

10°/sec, with a period of 1.1 seconds. The angular rate about the yaw axis

remained within 6° /sec at about the same period until 0 : 7-20, when 10°/sec

was also exceeded. Depletion of the fuel in the manual control system occurred

at approximately Oh:hj:0k. Correlation of the control -stick deflections with

the spacecraft angular rates indicates that at least 50 percent of the controlinputs opposed the oscillatory motion and that about 25 percent were approxi-mately 90° out of phase, thereby producing no net effect. The remaining25 percent apparently augmented the motion. Extrapolation beyond ±lO°/sec andintegration of the rate traces yields a maximum swing of ±5° for this period.

At 0 : 7:39, the astronaut switched to the auxiliary damping system.

Pitch and yaw oscillation rates decreased to within 2°/sec in a -secondperiod, while the roll rate decreased from -20

e/sec to a nominal -ll°/sec.

The depletion of automatic system fuel is estimated to have occurred atapproximately OU:48:30. At OU:48:32, the oscillations again began to increase

71





• ••• •

solenoid valves. However, these valves were thoroughly inspected and tested

after the flight, and no abnormal characteristics were disclosed. These

valves operated properly during tests in air and a vacuum, and a conclusive

explanation for the thruster failure cannot be given.

With the exception of the 1-pound yaw thrust chambers, all remaining

chambers in both the automatic and manual subsystems functioned properly

throughout the flight.

The fuel consumption for the orbital phase of the MA-6 mission is

listed in table IV, and the fuel consumption for the reentry phase is shown

in figure 30.

The manual proportional control system exhibited proper operation at

all times, with the possible exception of an instance in which it appeared

that the control handle may have retained a slight deflection from neutral

following a control maneuver. This condition appeared to have been cor-

rected by the astronaut with no difficulty.

Hydrogen peroxide feed-line temperatures in flight: Temperatures of

the hydrogen peroxide feed lines of both the automatic-system 1-pound roll

and the manual-system roll thrust chambers were measured during flight.

The feed-line temperatures of the manual-system roll thrusters increased

approximately 30° F from the prelaunch ambient temperature as a result of

aerodynamic heating during powered flight. At the completion of spacecraft

turnaround, the temperature of the feed lines to the manual-system roll

thrusters was stable at approximately 106° F.

Temperatures of the feed lines to the automatic-system 1-pound thrusters

were only slightly affected by the powered-flight phase, and instrumentation

indicated a total increase of less that 10° F from the prelaunch ambient

temperature from those at the completion of spacecraft turnaround. Maximum

temperature in orbit was approximately 120° F for the roll-thruster feel

lines of the automatic system and approximately 106° F for the roll-thruster

feed lines of the manual system. The effect of solar radiation on feed-

line temperatures was evident from data taken during the second and third

passes. During sunlight periods, slightly increasing temperature trends

were noted but did not continue in periods of darkness.

The reentry heating effect on all instrumented feed lines was very pro-

nounced: The temperatures of the feed lines leading to the roll-thrusters of

the automatic and manual systems reached maximums of approximately 1 5° F and200° F, respectively.

Postflight inspection: All thrust-chamber assemblies in spacecraft 13

were disassembled 6 days after the flight for the purpose of visual examin-

ation and photography. This inspection revealed loose foreign particles

upstream of the fuel-metering orifices of both 1-pound yaw thrusters. These

particles may have contributed to the intermittent malfunction of these

thrusters. These foreign particles have been visually identified as

portions of fuel-distribution (Dutch weave) screens located downstream of the

73



. . • •• • ;•: ...: ..'

fuel-metering orifices. The time and method by which these particles went

upstream of the fuel orifice is unknown. Examination results are listed in

table V, and typical postflight photographs are shown in figure 31-

Environmental Control System

The environmental control system (ECS) provided adequate environmental

conditions for the astronaut throughout the flight. The astronaut reported

uncomfortably warm conditions after landing, and these conditions are discussed

subsequently.

System description.- The primary change in the ECS from spacecraft 9

(MA-5) was the addition of a constant -bleed orifice to the suit circuit.

This orifice intentionally provided a continuous oxygen flow of 750 cc/min,

which was greater than the astronaut's anticipated metabolic requirement.

The excess gas was exhausted into the spacecraft cabin.

Countdown. - The temperature of the main inverters increased to higher

than expected levels during the countdown. These increased temperatures

indicated that the Preon flow to the inverter cold plates , though adequate

during precount checks, was inadequate during the final count. Temperatures

of the 150 v-amp and 250 v-amp inverters at lift-off were l62° F and 120° F,

respectively.

Launch phase.- The launch phase was normal. The cabin and suit pressures

were maintained at a differential pressure of 5-5 psi above ambient during

ascent and were held at 5-7 and 5.8 psia, respectively.

Orbital phase.- Cabin and suit pressures were maintained at 5-75.8 psia, respectively, throughout the orbital flight. The decay in these

pressures that had been observed in previous flights was absent for three

primary reasons :

1. The low cabin leakage (less than 500 cc/min)

2. Excess oxygen exhausted into the cabin from the suit-circuit

constant-bleed orifice

3. A small amount of leakage from the secondary oxygen supply

The oxygen partial pressure agreed with suit pressure to within 0.5 psia

although it was consistently lower. A partial cause of this difference is

the presence of water vapor in the suit circuit, which added a partial pressure

of approximately 0.3 psi that is not included in the oxygen partial pressure

measurement. A more careful calibration than those made for previous flights

has resulted in the more satisfactory performance of this instrument.

The cabin air temperature, after the initial heating period, fluctuated

as expected when the spacecraft passed through the alternate periods of



•••

• ••••• • • •

darkness and sunlight. The astronaut reported that at least five attempts

to reduce cabin air temperature by increasing water flow to the cabin heat

exchanger resulted in illumination of the. excess water light. This light

indicated that the cabin heat exchanger was operating near its maximum ca-

pacity for the existing conditions. Nevertheless, the mean cabin air temper-

ature was steadily reduced after the first hour in oribt.

The suit-inlet temperature (fig. 32(a))varied between 65° and 75° F

during the orbital phase. The astronaut reported a coolant flow of

1.7 pound/hour to the suit heat exchanger, and a steam exhaust temperature

of 60° F. These values are both higher than anticipated and contradict each

other, since freezing of the exchanger would be expected at this flow rate.

The reported flow rate was validated by postflight tests, and the logical ex-

planation for the high steam-exhaust temperature is that there was an error

in instrumentation.

The 150- and 250-v-amp inverter temperatures (fig. 32(b)) increased

steadily from the launch values of 162° F and 120° F, to maximum values of

212° F and about 200° F, respectively. Postflight testing revealed that the

check valve between the coolant supply and cold plates was stuck in the closed

position and would not permit the coolant to flow to the cold plates in orbit.

The coolant tank was charged with 25 pounds of water before the flight. The

coolant quantity indicating system showed a usage of 7-2 pounds. Postflight

weighing indicated a usage of 11.8 pounds. The difference in calibration and

final system temperatures can account for about 3-8 pounds of the U.6-pound

discrepancy, while the remaining 0.8 pound is considered to be instrument

error.

Reentry and postlanding.- The maximum cabin temperature during the re-

entry and postlanding period was 103° F, which is undesirable but satisfactory.The suit-inlet temperature increased to 87° F during the postlanding phase.

This value is reasonable since the air temperature in the landing area was

76° F, and the suit compressor raises the temperature in the suit circuit by

approximately 10° F.

Postflight investigation.- Examination of the flight data and post-

flight checks of the environmental control system have revealed several

anomalies. As shown in figure 33 > "the secondary oxygen supply exhibited

an unexpected decay in pressure, which was first noted at approximately

01:50:00. However, it is not known when this decay began, since the

secondary oxygen bottle was serviced to about 8,000 psig before flight,

and the pressure transducer had a maximum indicating value of only

7,500 psig. For an unknown reason, the decay rate of the secondary supply

was essentially zero during the last three quarters of an hour in orbit.

Postflight tests indicate a damaged "0"-ring seal as the cause of the

pressure decay in the secondary oxygen supply. In contrast, the decay

rate of the primary supply was much lower than expected during the final

portion of the mission.

75



- . . :•: ..: ••£

Communications Systems

The communications system group includes the two ultra-high frequency(UHF) voice transceivers, the high frequency (HF) voice transceiver, tworadar beacons, two command receivers, and recovery communications equipment.The recovery equipment comprises two UHF direction-finding (DF) beacons, an

HF (DF) beacon, and an HF transceiver. Since a more accurate measurementof the performance of these systems can be made from ground-based stations,their performance is discussed and appropriate signal strength and coveragedata presented in the Mercury Network Performance section. For this reason,only a brief discussion of their effectiveness for the MA-6 mission is givenin the following paragraphs.

Voice system.- The voice system, used for two-way voice conversationsbetween the ground and spacecraft, is made up of HF and UHF transceiver units.From previous orbital experience and ground tests, it was known that theHF system had somewhat poorer voice fidelity but longer range than the UHFsystem. The UHF system, because of its slightly better voice quality, was

considered to be the primary system. From previous experience, it was knovnthat the range of the UHF system was approximately equal to the line-of-sightrange and was entirely adequate for a normal mission. The main voice trafficwas therefore conducted on the UHF system, with a small amount of trafficconducted on the HF system to verify system operation. Performance of thevoice system during the MA-6 mission was satisfactory.

Radar system.- Performance of the radar system, including the groundtracking and computing complex, was satisfactory and was such that the space-craft orbital trajectory was well defined by the end of the first orbitalpass. The system's continued tracking during the remaining passes resulted

in only minor changes to the orbital parameters already established.

Command system.- The onboard command system was exercised by trans-mission of 10 instrumentation-calibration signals from the ground during themission for calibration of onboard sensors and to obtain additional data onthe command-system inflight performance.

Recovery system.- Performance of the recovery system was entirelysatisfactory.

Electrical and Sequential Systems

Electrical system.- The spacecraft electrical system was of a specifi-cation configuration. One-ohm fuse resistors were installed in all squibfiring circuits. Inverter monitor lights, which indicated to the astronautwhen the 150-v-amp and 250-v-amp inverters were operating, were installed.Main bus voltage and d-c current were as expected throughout the mission.Fans and ASCS a-c bus voltages were also normal.

The inverter temperatures were about 162° F on the 150-v-amp inverterand 120° F on the 250-v-amp inverter at lift-off and increased graduallythroughout the mission to 212° F and about 200° F, respectively. This

76



increase indicated that the inverters received little or no inflight cooling.Inverter cooling is also discussed briefly in the Environmental ControlSystem section.

Sequential system.- The sequential system performed as was expectedduring the flight, with the following minor exceptions:

1. Spacecraft separation, periscope extend, and damping command to theASCS were received 3 seconds after the launch-vehicle telemetry in-dicated spacecraft separation. The time discrepancy is attributedto the spacecraft adapter not having electrically latched bolt-firerelays to keep the spacecraft separation circuitry armed duringposigrade ignition.

2. The heat-shield-deployed signal was prematurely indicated bytelemetry (segment 51), and it then cycled on and off several timesduring the flight. This signal was caused by a defective left-hand

limit switch. In addition to the mechanical failure of the limitswitch, the circuit design did not provide for optimum reliabilitywith the two redundant switches installed.

3. Because reentry was performed with the retropackage attached, the0.05g sequential event was prevented from occurring automatically.The 0.05g event had to be initiated by the astronaut. Retropackageseparation indications did not occur, since the electrical signal toinitiate the retropack separation was never given. Manual initiationof 0.05g also resulted in the cameras remaining on high speed fromretroattitude command until power was removed after landing.

k. The drogue parachute was deployed prematurely at about 2J,000 feet.The 21,000-foot barostats were removed after the flight, subjectedto a preinstallation acceptance test, and found to be functioningproperly. Further tests of the static system were performed andare discussed in the Mechanical Systems section of this report.

Instrumentation

The instrumentation system flown on the MA-6 mission was essentiallythe same as in the MA-5. Small changes between the two spacecraft were re-quired to accommodate those parameters associated with a manned mission.

Other changes are as follows:

1. Heat shield temperatures for MA-6 were measured by means of achromel-alumel thermocouple.

2. The MA-6 spacecraft contained only two cameras, the pilot observerand instrument observer. The earth-sky and periscope cameras weredeleted.

3. A different type of color film was used in the pilot-observer camera.

77



,. ... • •••• • • •• •• • ••• • • I> . ••• • •

4. For MA-6, the mixed sequence of events were superimposed on thevernier clock signal, which is used as a time reference and at apulse frequency of once per second.

A complete list of the instrumented parameters is included in table VI.

Telemetry.- Both telemetry transmitters exhibited a center-frequencyshift and signal-strength rise when the escape tower was jettisoned. Thiseffect was anticipated and resulted from a change in antenna voltage standingwave ratio. The change in voltage standing wave ratio is generally notsignificant enough to prevent the transmitters from being adjusted to bringthe center frequencies within specification for both tower-on and tower-offconditions. The signal strengths and deviation from center frequency, asread by AMR when the spacecraft passed over Cape Canaveral, are shown intable VII.

Data quality.- The quality of the data reduced from the onboard tapewas very good. Scatter and noise caused by tape speed variations were in-significant and easily compensable on the continuous channels. The onlyproblem encountered occurred during the time of maximum exit dynamic pressure,when vibrational effects on the recorder caused almost complete loss of dataat about ^0 seconds after lift-off. This scatter and noise persisted forapproximately 50 seconds. However, the telemetered data during the sameperiod covered the lapse.

A comparison of the hand-controller (stick position) data, as instru-mented on continuous and commutated channels, shows the definite advantageof instrumenting this parameter by means of a continuous channel. About

50 percent of the stick-position data is lost on the commutated channel, andthat which remains is difficult to interpret.

Photographic.- The instrument-observer camera malfunctioned during the60-second test at T-55 minutes in the countdown. The film had slipped outof the film gate, but since the camera continued to run, there was no indi-cation that it was operating improperly. The results obtained from thepilot-observer camera were satisfactory. The use of the Ectachrome ER colorfilm was an improvement over the types previously used.

Onboard timing.- The vernier clock channel malfunctioned throughout theentire flight. Each time the pilot-observer camera operated, a spuriouspulse was produced in the vernier clock signal. At times when the cameraoperated at high speed, the clock signal was rendered nearly useless.

Respiration sensor.- The most serious instrumentation problem during theMA-6 mission was in the respiration rate and depth channel. The circuitutilizes a thermistor sensor which is heated by a d-c voltage. The resis-tance of this sensor is changed when the pilot cools it by exhaling, and thevariation of resistance produces a change in voltage at the output of adirect-coupled transistor amplifier. The basic problem is that the thermistortemperature is also affected by environmental changes within the helmet.Because of these changes, the baseline of the respiratory signal

78 JOHFIDENTIAL



:':•;* :-.: .-..•• .•• : :« • - - - ™ . « ••• • » .. . .- - -

C OHFIDE3JTIAL *

varied from a value of 10 percent, which was set during the suiting procedure,

up to a maximum of 85 percent at lift-off. After lift-off, the baseline fell

steadily until it reached a low of 10 percent at 02:08:00. It then began to

rise again and had attained a level of 40 percent at loss of the telemetry

signal. The sensitivity of the signal was degraded in direct proportion to

the baseline shift, since the sensitivity decreases as the baseline increases.

The sensitivity problem is further complicated by the fact that the position

of the sensor is not fixed to the position of the pilot's head. In the

MA-6 mission, much of the data was lost because the pilot was not breathing

directly on the sensor. Both subcarrier oscillators which encoded the

respiration data performed within specification and did not contribute to

the above effects.

Fuel-quantity indicators.- The fuel-quantity indicator for the manual

system registered an average error of approximately 8 percent more fuel than

the actual amount for fuel quantities below 70 percent. The fuel-quantity

indicator for the automatic system registered an error of approximately

3 percent more than the actual amount for all fuel levels.

Heat Protection System

Performance of the MA-6spacecraft heat protection system during the

mission was satisfactory in all respects. Minor cracking was evident, as

expected, during the postflight inspection. This condition resulted from

the heating experienced during reentry, which is considered to be identical

to that of MA-5.

Since the heating data recorded during the MA-6reentry are not nearly

so comprehensive as those for MA-5, the reentry temperature survey for thelatter flight is presented in figure 3^. Temperature readings on the conical

and cylindrical portions of the spacecraft employed thermocouples spot welded

to the inboard side of the exterior shingles. The data presented are well

within the specified range and are therefore considered nominal. For a com-

parison of these results obtained theoretically and during preflight wind-

tunnel research, see reference 9-

Mechanical and Pyrotechnic Systems

All mechanical and pyrotechnic systems functioned normally with the

exception of the premature deployment of the drogue parachute and an indi-

cation of heat-shield release during the orbital flight phase. These

anomalies, along with general systems performance, are discussed in the

following paragraphs.

Parachutes.- The deployment and performance of the drogue and main

parachutes were satisfactory. Since neither parachute was recovered, a

detailed postlaunch visual inspection could not be made. However, observation

by the astronaut verified that both parachutes were undamaged during the de-

ployment phases and throughout descent. The main parachute deployed at a

pressure altitude of 10,000 feet, as determined from onboard measurements.

This is within specification limits of 10,000 ±750 feet.

79



The drogue parachute deployed at an altitude about 27,000 feet, whichis above normal. Onboard pressure measurements (commutated) indicate apressure altitude of approximately 29,000 feet at the time of drogue para-chute deployment, and the integrated trajectory is consistent with a drogueparachute deployment at about 27,000 feet. The astronaut reported an al-titude of 30,000 to 35,000 feet was indicated at the time of drogue para-

chute deployment. The drogue parachute barostats actuated properly withinspecification pressure altitudes of 21,000 ± 1,500 feet in postflight tests.

The possible causes of an early deployment of the drogue parachute which

are listed below were examined separately and in appropriate combinations.

1. Inertial actuation of the barostat as a result of spacecraft

oscillations

2. Dynamic pressure q effects caused by oscillations of the spacecraft

3 - Actual astronaut actuation of the manual drogue parachute switch

^ 4 - . Structural failure within the parachute compartment, particularlythe drogue parachute mortar cover

5. Static pressure buildup in the parachute compartment

6. Spurious or random signals in the drogue parachute mortar-firing

circuit

The first source was ruled out after testing, since the 2g accelerationimposed by the spacecraft oscillations are within the rigid qualification

specifications governing the flight hardware. Strong evidence exists thata dynamic pressure could not have built up without being sensed by the staticpressure transducer in the parachute compartment. Item 3 was eliminated asa cause through postflight interviews with the astronaut and examination ofphotographs taken in the spacecraft during the period in question. The pilot

stated that he had raised his arm about halfway to the switch when automaticdeployment occurred. Postflight examination of the parachute compartmentrevealed no structural failure, and recorded sequence data indicate that themortar was fired by a nominal electrical signal, disproving item U. Tests atMcDonnell Aircraft Corporation have shown that an increase in the amplitudeof pressure oscillation present in the parachute compartment is sensed by the

pressure transducer; therefore, this increase in pressure would have beenevident on the flight record. The last item represents the most plausiblecause since the data reveal that a release signal was apparently receivedby the drogue parachute mortar.

Postflight tests have shown the entire circuitry, including the barostats,to be sound; however, these tests were conducted under static conditions and,therefore, are inconclusive. The definite cause of the premature drogue para-chute deployment is unknown.

80



• ••. ; • • •i i i . .

Rockets and pyrotechnics. - A postflight examination of the spacecraftand an analysis of the pertinent data indicate that all rockets and pyro-technics functioned as intended. It cannot be determined whether certainpyrotechnics actually ignited (such as redundant clamp-ring bolts andtower-jettison rocket ignition), since the available evidence shows onlythat the resulting function was satisfactory.

Explosive -actuated hatch. - After the spacecraft was secured on boardthe recovery ship, the astronaut initiated the hatch explosive mechanismthrough the use of the internal actuator on the hatch. The hatch mechanismappeared to have operated satisfactorily.

Landing -shock attenuation system. - The landing -shock attenuationsystem performed normally, as evidenced by the astronaut's statements andfrom postf light examinations. The landing bag was found to be torn inseveral places (see also Postflight Inspection section), but no restrainingcables or straps were broken. The usual minor damage to ablation shield re-

taining studs and to the bulkhead-protective shield was experienced as aresult cf recontact of the heat shield. The main pressure bulkhead was un-

damaged except for a small dent near the center.

A comprehensive postflight inspection of the limit-switch circuitryand hardware revealed that both lateral and axial translation would generatea signal, altho.ugh the switch is designed to actuate only after a rotationaldeflection. The cause of this improper operation was traced, after dis-assembly, to a poorly machined shaft. (See fig. 35-) The shaft was dentedin several places and bent some 5° or 10° . The pilot reported noises andother indications of heat-shield deployment when he manually initiated thissequence soon after main parachute deployment; thus the conclusion may be

drawn that the shield mechanism was not unlatched in orbit.

Postflight Inspection

The general condition of the spacecraft upon completion of the missionwas excellent, as indicated in figure 36- The exterior of the spacecraftshowed the usual slight discoloration caused by aerodynamic heating. Therewere also deposits of 2024 aluminum alloy which had evidently been trans-ported in a molten state and had adhered to the surface of several widelyseparated shingles. The aluminum retropackage which was retained into re-entry was the source of these deposits. A brownish deposit was found on a

portion of the spacecraft window exterior surface. As on previous flights,the window was found to be fogged with water condensate between the twoouter panes .

Structure. - The spacecraft did not experience any structural damagewhich would have compromised the safety of the mission in any way.

Ablation shield. - The external surface of the shield (see fig. 37)had charred in the normal pattern. The edge of the heat-shield center plughad separated as expected and extended outward approximately 0.5 inch.However, it remained attached at the center. A circular sector of the

81





•. :• • •• • . ::••: :•• ••iigP":-. :•: .1- .:• •:" •:;. :.:

TABLE IV.- FUEL CONSUMPTION

Fpata accuracy to ±0.5 pound of fuell

Flight phase

Launch

Turnaround

and damping

Orbital pass 1

Orbital pass 2

Orbital pass 3(to retro-sequence)

Retrosequence to

0.05g

0.05g to drogue

parachutedeployment

Drogue parachutedeployment tomain parachutedeployment

Automatic system

Fuel used,Ib

0

5 - 8

k.2

6.0

8.6

k . o

a?.

0

Fuel remain-

ing, Ib

36.0

30.2

26.0

20.0

11A

7

0

0

Manual system

Fuel used,

Ib

0

0

0.6

11.8

5.2

5 - 6

al.2

0

Fuel remain-

ing, Ib

2k.k

2k. k

23.8

12.0

6.8

1.2

0

0

Fuel depletion occurred during this period

83



TABLE V.- RESULTS OF POSTFLIGHT EXAMINATION OF THRUST CHAMBERS

Thrust-

chamberassembly

1-pound,yaw-right

1-pound,

yaw-left

1-pound,pitch- down

1-pound,pitch-up

1-pound,counter-clockwiseroll

1-pound,clockwiseroll

Automatic6-pound,clockwiseroll

Automatic6-pound,counter-clockwiseroll

Heat-

barrierscreen

Clear

Clear

Clear

Clear

Clear

Clear

Clear

Clear

Orifice

condition

Five particleson upstreamface

Numerous dust

size particleson upstreamface

One particleon upstreamface

Nine particleson upstream

face

Clean

Clean

Clean

Clean

Dutch-weave

condition

Top screenburned andheavilyeroded incenter

Top screen

burned andmoderatelyeroded

Small holein topscreen andslightlyburned

Top screenburned and

practicallyeroded incenter

Normal-onlyslightlydiscolored

Normal-onlyslightly

discolored

Normal

Slighterosion

Remarks

Particle diameterapprox. same as0.016 inchorifice andlength somewhatlarger than0.016 inch

Particle size of

the order 100 to200 microns

Particle sizeof the order100 to 200microns

Particle sizesame as those in

1-pound,yaw -right,above



'CABLE V.- RESULTS OF POSTFLIGHT EXAMINATIONOF THRUST CHAMBERS - Continued

Thrust-chamberassembly

Automatic24-pound,

pitch-up

Manual2k- pound,pitch-up

Automatic24-pound,

pitch- down

Manual24-pound,

pitch- down

Automatic24-pound,yaw-left

Manual24-pound,

yaw-left

Automatic24-pound,

yaw-right

Manual

24-pound,

yaw-right

Heat-barrierscreen

Clear

Clear

Clear

Clear

Clear -appearsas only1 screen

Clear

Clear

Clearand wet

Orificecondition

Clean

Clean

Clean

Clean and wet

Clean

Clean

Clean

Clean and wet

Dutch-weavecondition

Not applicable

Not applicable

Not applicable

Not applicable

Not applicable

Not applicable

Not applicable

Not applicable



TABLE V.- RESULTS OF POSTILIGHT EXAMINATIONOF THRUST CHAMBERS - Concluded

Thrust-

chamberassembly

Manual6-pound,counterclock-wise roll

Manual6-pound,

clockwiseroll

Heat-barrierscreen

Clearand wet

Clear

Orifice

condition

Clean

Clean

Dutch-weavecondition

Normal

Normal

86



* •• •<

•• ••••• • ••

• <••

TABLE VI. - ISSTHUMENTED PARAMETERS TOR MA-6

(a) Commutated quantities

High-frequency

channel

1

2

3

k

5

6

7

8

9

10

11

12

1}

I k

15

16

17

18

19

20

21

22

25

2"t

25

26

27

28

29

30

31

32

33

low-frequencychannel

1

2

3

It

5

6

7

8

9

10

11

12

13

lit

15

16

17

18

19

2 0

21

22

23

2l*

25

26

27

28

29

30

31

32

33

Parameter

Zero ground reference, volte

a-c amplifier power supply monitor, volts

"B" c o m m a n d receiver signal strength, nv

Suit pressure, pflia

Cabin air temperature, *P

Cabin air temperature, *F

Z-axie acceleration, g units

Pitch attitude A S C S calibrator, deg

Roll attitude ASCS calibrator, deg

Yaw attitude A S C S calibrator, deg

Low rollclockwise manual-fuel-llne temperatures, "F . .

Low roll counterclockwise manual -fuel-line

Low roll counterclockwise automatic-fuel-line

Low roll clockwise automatic-fuel-linetemperature, °f

Static pressure, psia

Stick position, roll, deg

Stick position, pitch, deg

Stick position, yaw,deg

Elapsed time (0 to 1 minute), percent

Elapsed time (0 to 10 minutes ), percent

Elapsed time (0 to 10 hours), percent

Instrument range,

0- to 100-percent full scale,

unless otherwise noted

o

full scale)0 to 80 (20- to 68-percent

full scale)

-0.2, 7.1*-, 15 3 (0- 98- 0-percent

full scale)

-100 to 7 600

full scale)

-0.415 to 0 375(15-to 80-percent

full scale)

full scale)

-Jl to 35

-120 to 17

-130 to 190

-12 to 260

-11 to 239

full scale)

full scale)

full scale)

full scale)

0 to 100

0 to 100

87



TABI£ VI. - CBTRUMEHTED PARAMETERS POP MA-6 - Continued

(a) Connuutated quantities - Continued

High-frequency

channel

^4

^57

38

59

40

41

42

43

kk

45

46

47

U S

49

50

51

52

55

5

55

56

57

58

59

60

61

62

65

64

65

66

67

68

69

70

71

72

low-frequencychannel

^4

55

56

57

38

59

40

41

42

4 5

44

45

46

47

48

49

50

51

52

55

54

55

56

57

58

59

60

61

62

65

64

65

66

67

68

69

70

71

72

Parameter

Mayday

Calibration, Z-cal, normal R-cal (also 25-, 50->

High-pressure reaction Jet solenoids (- pitch) . . . .

Instrument range,



0 to 100

0 to 100

0 to 100

On -off

500 to 2,200 (4- to 95-percent

full scale)

600 to 2 200 (6 5- to 89^>ercent

full scale)

95 to 120 (68.5- to 89.3-percent

full scale)

0 to 50

-0.415 to 0.58 (19- to 80-percent

full scale)

-4 0 to 4 9

-0.415 to 0.575 (15- to 80-percent

full scale)-5 2 to 5 5 (15- to 85-percent

full scale)

-51 to 55

O n -off

O n -off

O n -off

0 to 80 (30- to 78-percent

full scale)


full scale)

On -off

On -off

0 to 565

On -off

On -off

On -off

On -off

On -off

On -off

On -off

On -off

On -off

On -off

On -off

On -off

On -off

On -off

15.5 to 25.4

15.2 to 25

On -off

On -off

On-off


full scale)


full scale)

On-off



TABLE VI. - INSTRUMENTED PARAMETERS FOR MA-6 - Continued

(a) Coonnutated quantities - Concluded

High-frequency

channel

Low-frequency

channel

Instrument range,



73

75

76

77

78

79

80

e i

8 2

93

81*

87

90

73

71*

75

76

77

78

79

80

81

82

83

8*

87

90

Y-axls acceleration, g unite

Y-axls acceleration, g units

X-axis acceleration, g units

X-axis acceleration, g units

Z-axls acceleration, g units

Heat shield temperature (thermocouple), *F

ASCS bus voltage, v a-c

High-pressure reaction Jet solenoids (-yaw) ....

High-pressure reaction Jet solenoids (+yaw)

Retrorocket temperature, °F

High-frequency telemetry transmitter temperature, *F

low-frequency telemetry transmitter temperature, *F

Cabin pressure, psla

d-c voltage monitor, volts

Coolant pressure, pslg

Horizon scanner roll ignore

Roll horizon scanner output monitor, deg

0. 05g relay actuation

Pitch horizon scanner output monitor, deg

Synchronize pulse

Synchronize pulse

-0.1*15 to 0.38 (15- to 80-percent

full scale)

-k. 0 to It. 9

O.ltl5 to 0.375 (15- to 80-percent

full scale)

-3.2 to 3.5 (15- to 85-percent

full scale)

-31 to 35

-lUO to 2,^70

90 to 125 (6 .7- to g't-percent

full scale)

On-off

On-off

-16 to lUO

8 to 320

-16 to 326

-0.2, 7-1*, 15.2 (0-, 101-, and

0-percent full scale)

18 to 28

213 to USg

On-off

-37.5 to 33

On-off

-53 to 18.5

89



TAB1£ VI. - IlBTRUMEirrZD P A R A M E T E R S TOR MA-6

(b ) Continuousquantities

- Concluded

R-equency

n *A

1 3

x o

10.5

Frequency

0 •&

0 73

1 7

5 *

H igh-frequency system

High-frequency conmutator, pulse amplitude modulation

low-frequency system

I n s tr u m e n t range,& - to 100-percent full scale.


-9.9 to 10

^5 5 to 31 5

12-9 to 30

±12(22- to 79-p*rcent full scale)

Instrument range,0- to 100-percent full scale


-10.3 to10.5

±10.5

-0.6 to50.6

-10.5 to 12

(c) Onboard tape recorder track assignment

Trade Information

Open

High-frequency telemetry multiplex

V o ic e

Pulse damped modulation, high frequency

Pulse damped modulation, low frequency

Low-frequency telemetry multiplex

Open

90 DNFIDENTIAL





> • • 1 •

200

Figure 29.- Example of the 180° yaw maneuver.

92



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105



• • • • •

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Figure 36.- Postflight photograph of spacecraft,

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107



AEROMEDICAL ANALYSIS

The aeromedlcal investigations conducted for the MA.-6 mission aredivided into two areas: (l) the preflight and postflight clinical exami-nations (static examinations) and (2) the preflight and inflight physio-logical studies (dynamic studies). These investigations are designed toascertain the state of the astronaut's health and to provide informationreflecting human responses to space flight. The MA.-6 mission provided a

period of weightlessness of sufficient duration (k hours) for the pilot's

physiological responses to attain a relatively steady state, in contrast tothe shorter Mercury-Redstone flights (refs. 10 and 11), where little timewas available in a zero-gravity state for the astronaut's physiologicaladjustment mechanisms to stabilize.

The astronaut's activities during the time immediately prior to hisinsertion into the spacecraft have some effect on his countdown and flightresponses. For this reason, his activities for the approximate 8-hourperiod prior to his arrival for insertion into the spacecraft are listedin table VIII.

The pilot began a 72-hour, prelaunch low- residue diet on February l6, 1962.On the night prior to flight, the pilot obtained k hours and 50 minutes ofdozing, light sleep. No medication was administered.

Clinical Studies

Detailed medical examinations were performed prior to the MA.-6 flight,and Ri.Tn-na.-r investigations were conducted as soon after the flight as re-covery practices permitted. Initial examinations were made to determine theastronaut's state of health and his medical fitness for flight. In addition,such clinical evaluations serve as baseline medical data which may be latercorrelated with inflight physiological information.

The sources of clinical medical data regarding the astronaut are listedbelow:

1. Prior physical examinations beginning with astronaut selection

in 1959

2. Detailed preflight clinical examinations conducted onJanuary 22, 1962, and February 12, 1962

3. Preflight examination conducted on the morning of launch

4. Postflight medical evaluations on the recovery ships and atthe Grand Turk Island Medical Facility

108



• • •• •• •••

• • •

The numerous preflight examinations of the astronaut disclosed no sig-nificant medical abnormalities; his physical and mental health remained

excellent throughout the preflight period.

The postflight examination of the astronaut began with his emergence

from the spacecraft on board the recovery ship some 39 minutes after landing.The pilot was described as appearing hot, perspiring profusely, and fatigued.He was lucid, although not talkative, and voiced no medical complaints other

than being hot.

After his pressure suit had been removed and the astronaut had showered,

he became more communicative and described a mild sensation of "stomach un-easiness" or "stomach awareness," which occurred during the 17 minutes on thewater while awaiting recovery. This sensation did not commence until after

spacecraft landing and cleared spontaneously within 1 hours. Neither nausea

nor vomiting was experienced. This stomach uneasiness can be attributed toseveral factors. First, there is the combination of temperature and humidity

immediately after landing. The suit inlet temperature was 85° F; the cabinair temperature was 103° F. The ambient air temperature was 76° F, relativehumidity was 60 to 65 percent, and the sea-water temperature was 8l° F. Asecond major factor contributing to this gastrointestinal condition is themoderate dehydration of the astronaut. This dehydration is evidenced by a

weight loss of 5r^ pounds during the mission, diminished urine output with

increased specific gravity for the 2i)--hour postrecovery period, increased

blood concentration, and the recovery physician's clinical impression.

The astronaut had a minimal fluid intake during the 13 hours from break-

fast at 2:50 a.m. e.s.t. to shipboard at 3: 5 P-m. e.s.t.j during this periodthe equivalent of only 9^ cc of water was ingested as applesauce puree. Theonly other oral intake during the flight was one 5.0-gram sugar (xylose) tabletused as a test of intestinal absorption; the results of this test were normal

(table IX). His urine output during this period was 800 cc which he voided ina single specimen just prior to reentry. (See table X.) In examining weightloss, the pilot had the fluid intake and output shown in the following table:

Countdown

Inflight

Postflight, ship

Total

Urineoutput, cc

0

a800

0

800

e. s. t.

_ _ _ «

2:00 p.m.

""•""••"

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0

V265 (iced tea)

2 0 (water)125 (coffee)J2k

e. s. t.

_ . _ _ _

1 1 : 1 4-8 a.m.

3:^5 a.m.6:30 p.m.

6:50 n.m.

rj

Specific gravity, 1.016

119.5 grams of applesauce puree (78.7-percent water)

109





The countdown period provided baseline preflight information.comparative measurements were available from the Mercury-Atlas three-

simulation, the pad simulated launch, simulated flights,the launch attempt of January 27, 1962. Environmental control system datacorrelated with physiological responses where appropriate.

Bioinstrumentation.- The biological sensors used for the MA-6 mission were:leads, respiratory rate sensor, and body temperature sensor. A blood-

suring system (BPMS)was utilized in flight. The BPMS consisted ofpneumatic nylon cuff placed on the left upper arm and a microphone located

over the brachial artery. The cuff was inflated manually tothe blood-pressure reading. The blood-pressure record consisted of the

pulses superimposed on a cuff-pressure decay curve. This recorddisplayed on the second EGG channel. Comparison of preflight and postflight

of the blood-pressure system showed no significant change.

The total biosensor monitoring time, from astronaut insertion until justto landing, was 8 hours and 33 minutes. The biosensor readout quality

excellent throughout the countdown and flight, with the exception of theiratory trace. As in prior manned flights, variation with head position

density combined to reduce the quality of this trace. There were briefnoise on both EGG channels during countdown and flight, usually

ccurring during vigorous pilot activity.

Preflight.- Figure 38 depicts the pulse rate, respiration rate, bodytem-and blood-pressure values recorded during the

countdown. Times at which significant events occurred are indicated athe bottom of the figure. Values for the physiological functions obtained from

he simulated launch of January 19 and the launch attempt of January 27, 19 2,also shown.

Heart and respiration rates were determined by counting the rates forseconds every 3 minutes until 10 minutes prior to lift-off; thereafter,

0-second duration counts were made each minute. During approximatelyminutes in the transfer van, the astronaut's heart rate varied from

ts per minute, with a mean of 72. His blood pressure was™n Hg. The heart rates during the canceled-flight countdown of27 varied from 60 to 88 beats per minute, with a mean of 70. These

the same as those observed during the MA-6countdownthe mean heart rate was 68 beats per minute. Respiration rates were20 breaths per minute. Blood-pressure values from the simulated launchapproximated those observed during the MA-6countdown. A pulse rate of

10 beats per minute and a blood pressure of 139/88 mm Hg were observed duringcountdown prior to lift-off. The low suit-inlet temperature maintained duringountdown resulted in the pilot's feeling cold and was accompanied by a fall in

temperature from 98.6° F at insertion into the spacecraft to 97-6° at

An examination of the electrocardiographic waveform obtained during thecountdown revealed a number of variations in the pacemaker (thepoint at

ulus of the heart beat originates) activity. These variations

111



•• •••• • •• • »•

included sinus pauses, sinus bradycardia (slowing), premature atrial andnodal beats, and premature ventricular beats. On several instances, some ofthese reported findings occurred with deep respiration. Similar findingswere recorded from the simulated launch of January 19 and the canceled flightof January 2J. In addition, a brief run (l6 beats) of atrial rhythm with arate of 100 beats per minute occurred during countdown, and an isolated run(19 beats) of rhythm originating adjacent to the atrio-ventricular node withaberrant conduction occurred during the attempted launch of January 27-However, these arrhythmias were not observed at any other time. They arethought to represent normal physiologic variations. Samples of MA.-6 recordstaken at the time of astronaut insertion and at T-50 seconds are shown infigures 39 and. -0, respectively.

Flight.- Figure ^1 depicts the inflight physiological data. Values fromthe Mercury-Atlas, three-orbit centrifuge simulation are included for com-parison. Heart rates in beats per minute were counted at 30-second inter-vals during the MA-6 launch and reentry phases and for 30 seconds at 3-minute

intervals throughout orbital flight. Because of the variation in respiratoryrecording quality, rates were counted for 30 seconds whenever quality permitted,and these rates varied from 8 to 19 breaths per minute throughout flight.

The heart rate from lift-off to spacecraft separation (powered-flightphase) reached a maximum of 110 beats per minute. The heart rate variedfrom 88 to Il4 beats per minute during the first 10 minutes of weightless-ness. With the exception of periods of planned, calibrated exercise, itremained relatively stable, with a mean rate of 86 beats per minute duringthe next 3 hours and 45 minutes of flight. At the time of retrorocket firing,the rate was 96 beats per minute. In the reentry-acceleration and parachute-descent portions of the flight, the highest pulse rate was 13^ beats perminute which occurred during the period of spacecraft oscillation duringreentry prior to drogue parachute deployment. This rate was the highest notedduring the mission. These rates suggest that acceleration, weightlessness,and return to gravity were within physiologically tolerable limits.

The EGG variations noted during the preflight observation period werenot observed in flight, and analysis of the inflight record revealed onlynormal sinus rhythm with short periods of sinus bradycardia and sinusarrythmia. There were rare periods when trace quality deteriorated to apoint that only the heart rate could be determined. Variations in EGGtracings noted during the pilot's Mercury-Atlas, three-orbit centrifuge

simulation included sinus arrythmia, sinus bradycardia, atrial and nodalpremature beats, and rare premature ventricular contractions. These areinterpreted as being normal physiological variations.

Ten blood-pressure determinations were made in flight, the firstat 00:18:30 and the last at 03:1 :00. Values, as shown in figure 4l,

range from 119 to 1 -3 mm Hg systolic and from 60 to 8l mm Hg diastolic.

112



: *

The mean blood-pressure values from various other blood-pressure data sourcesare presented in the following table:

Data sources

Physical exams

Procedures trainer

3-orbit Mercury-Atlas centrifugesimulation

Launch- pad tests

MA-6 countdown

MA-6 flight

Number ofdetermi-nations

14

15

56

26

14

10

Mean bloodpressure,

mm Hg

110/66

121/76

114/80

104/76

123/87

129/70

Mean

pulsepress-

ure,i r o n Hg

44

45

34

28

36

59

Systolicrange,

mm Hg

98 to 128

110 to 132

92 to 136

91 to 125

101 to 139

119 to 143

Diastolicrange,

mm Hg

60 to 80

66 to 87

68 to 92

64 to 91

83 to 93

60 to 8l

The mean pulse pressure during the MA-6 flight shows a slight wideningwhen compared with preflight values taken in the centrifuge. The meanblood-pressure value from procedures trainer simulations was 121/76. Thewidened pulse pressure, which appeared after 1 hour of flight, is of un-certain physiological significance. Samples of physiological data from theonboard record are shown in figures 42 to 44.

A photograph of the inflight exercise device is shown in figure 45.Exercise was accomplished by a series of pulls on the elastic bungee cords.Exercise over Zanzibar during the first orbital pass elevated the pilot'spulse rate from 80 to 134 beats per minute after 30 seconds. The heart ratereturned to 84 beats per minute within 2 minutes. The blood pressure beforeexercise was 129/76 mm Hg and 129/74 mm Hg after exercise, which is similarto preflight trainer readings.

The environmental control system effectively supported the pilotthroughout the mission. It should be noted, however, that body temperature

gradually rose from a lift-off value of 97-6° F to 99-5° F at biosensordisconnect. The suit-inlet temperature increased slowly during most of theflight, with a more rapid rise after reentry and during parachute descent.

During the period immediately following descent, the suit-inlet temperatureincreased approximately 1° F per minute for a 15-minute period and probablycontributed to the pilot's overheated status observed at egress. Since bio-

sensor disconnect occurred 13 minutes before loss of signal, the maximumbody temperature may not have been observed.

113



•• ••• • ••• • •• .. .•• •• •• • . . • ••» ••• • •• t "I J .' • I • • . • • •

* •• •• • • . , . .

Pilot Inflight observations.- The astronaut's voice reports wereconsistently accurate, confident, and coherent through all phases of theflight. His voice quality conveyed a sense of continued well-being. Hismental state appeared entirely appropriate for the situation. The pilot'smood and level of performance was satisfactorily conveyed by his voice

reports. His prompt responses to ground transmissions and to sounds fromthe spacecraft suggest no apparent decrement in hearing ability. Visualacuity was maintained, and his report of visual perceptions, especiallywith regard to colors, was confirmed as accurate by the inflight photographs.The pilot's voice report contained observations of physiological significance.During his postflight debriefing these reports were amplified. Those con-sidered most significant are discussed in the following paragraphs.

No disturbances in spatial orientation were reported, nor were anysymptoms suggestive of vestibular (inner ear) disturbances described duringthe flight. Voluntary, rapid head-turning movements produced no unpleasant

sensations. No sensory deprivation or "break-off phenomena" were noted.

Immediately after sustainer engine cutoff (SECO), the astronautexperienced a brief sensation of tumbling forward, similar to thatdescribed by the astronauts in the suborbital missions. This sensationended promptly with no accompanying nausea. Coincident with retrorocketignition, a feeling of movement opposite from flight direction was noted.This feeling could be expected with the sudden change in spacecraftvelocity. The pilot noted no difference in the sensations associatedwith reentry accelerations from those experienced during launch.

Food (applesauce puree and a xylose tablet) chewing and swallowing were

accomplished without difficulty. No liquid as such was ingested during flight.

The pilot urinated without difficulty shortly before reentry. Hedescribed "normal" sensations of bladder fullness with the associated urgeto urinate.

The astronaut described weightlessness as a "pleasant" sensation.Control manipulation was not a problem, and there was no observable per-formance decrement. The restraint harness and couch combination was re-ported to be comfortable.

In summary, the MA.-6 mission provided a period of extended weight-lessness during which the astronaut's physiological responses apparentlystabilized. The values attained were within ranges compatible with normalfunction. No subjective abnormalities were reported by the pilot.

llU



TABLE VIII.- AEROMEDICAL EVENTS PRIOR TO LAUNCH

Date e.s.t. Event

February 19, 1962

February 20, 1962

9:30 p.m.

2:20 a.m.

2:50 a.m.

3:05 a.m.

4:28 a.m.

5:05 a.m.

5:20 a.m.

5:58 a.m.

6:06 a.m.

6:25 a.m.

9: 7 a.m.

Retired

Awakened and showered

Breakfast

Physical examination

Suiting started

Entered transfer van

Arrived at launch pad and remainedin transfer van

Ascended gantry

Inserted into spacecraft

Countdown resumed

Launch

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(a) Countdown, 06:00 to 08:00 e.s.t.

Figure J > Q . - Respiration rate, pulse rate, body temperature, suit-inlet

temperature, and "blood pressure measured during countdown.

122



>• •<• • •

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(b) Countdown, 08:00 to 09:^7 e.s.t. (lift-off).

Figure 38.- Concluded.

123



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130



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• •• • •

ASTRONAUT FLIGHT ACTIVITIES

A brief review of the astronaut's activities in preparation for andduring the MA-6 mission is presented in this section.

Preflight Training Si.nrmia.Ty

Spacecraft checkout activities.- The astronaut's participation in pre-flight checkout and testing provided him vith the opportunity to becomefamiliar with the MA-6 spacecraft and launch-vehicle systems. His activitiesduring this period are summarized in table XV. He spent 25 hours and55 minutes in the spacecraft itself; and many more hours were consumed beforeand after each launch-pad checkout operation in preparation, trouble-shooting,observation, and discussion.

Training activities.- Table XVI summarizes the astronaut's trainingactivities in the procedures trainer. During the period from December 13, 1961,to February 17, 19^2, the astronaut spent 59 hours and ^5 minutes accomplishing70 simulated missions and experiencing 189 simulated systems failures. Themain emphasis during this period was on procedural training, particularlylaunch and early mission failures usually calling for a mission abort. Theastronaut accomplished several three-pass missions on the trainer, as well asseveral simulations involving the Mercury Control Center (MCC), the blockhouse,and the entire Mercury Network. The astronaut also spent appreciable time inbriefings on the various subsystems, flight planning, and in individual study.

Flight preparedness.- The astronaut achieved a high level of skill and

knowledge several weeks prior to the actual launch date. There was a gradualreduction in the intensity of the preflight training program, particularlyon the procedures trainer. Nevertheless, there was no decline in his levelof preparedness.

The astronaut's comments during postflight debriefing sessions indicatethat he acknowledges the value of his preflight training. His major comments

were: (l) that his participation in the spacecraft checkouts was very useful,(2) that personal briefings from systems specialists were helpful, and(3) that the air-lubricated-free-attitude (ALFA) trainer was less valuablethan the procedures trainer as an attitude control trainer.

Chronology of Pilot's Activities During Flight

Figure k6 is a simplified chronology of the pilot's activities duringthe MA-6 flight from lift-off to landing. Identification of continentallimits, certain celestial observations, photography, and communications modesare not included in the table, since the communications tapes do not providesufficient time correlation. Spacecraft systems problems during the flightprevented the pilot from completing all of the planned tasks.

133





••• ;•.: :.: '..\ • :• : r. :.:

.. *" * ~ v* * * * * ** **

Retrofire control.- The astronaut backed up the ASCS during the retro-sequence and retrofire events by using the manual proportional control mode.It is impossible to assess individually the operations of the ASCS or ofthe astronaut at this time. The attitudes did not deviate more than ±3°during this event.

Reentry pitch maneuver.- The manual proportional control system and therate and attitude indicators were used by the astronaut in pitching up toreentry attitude. As can be seen from figure 50 he performed this maneuverwith precision and was well within the capability demonstrated in groundtrainers.

Reentry damping.- The early part of the reentry through peak accelerationwas controlled by employing the fly-by-wire and manual proportional controlsystems, and the oscillations during this period were small. (See fig. 51-)After c4:V7:00, the pitch and yaw oscillations increased rapidly, and thepilot's control inputs did not appreciably reduce these rates. It is evident

from an analysis of the fuel usage that his manual fuel was depleted atC4: l j - 7 ; Oh. The fly-by-wire .control mode apparently still had fuel available,as indicated by the subsequent effectiveness of the auxiliary damping modeat 0 :47: 2. The ultimate lack of satisfactory control is attributed to thechange in control effectiveness because of the early depletion of manualcontrol system fuel.

Pilot's use of external reference.- The pilot reported that, by usingexternal reference, he was generally able to orient the spacecraft withoutdifficulty, particularly about the roll and pitch axes. The pilot statedthat there was a period of learning associated with use of the periscopeand the window as yaw references, and he felt that by the end of the flight

he was able, on the daylight side, to adjust yaw within a few degrees.During the flight, he developed the procedure of pitching down to -60° inorder to increase the apparent terrain drift in the window caused by orbitalvelocity. At a pitch attitude of -3^-°, the apparent velocity of terrainmovement is 0.60°/sec, whereas, at -60° pitch attitude, the apparent velocityis increased to 1.4l°/sec. The greater apparent drift as the vehicle pitchestoward the nadir point aids in the determination of yaw attitude by increasingthe ratio of the terrain movement resulting from orbital velocity to theapparent terrain movement resulting from a spacecraft attitude rate in yaw.The pilot preferred the window over the periscope for yaw alinement. Hedisliked the periscope high-magnification view because of the unclear area

at the border between the high- and low-magnification views.

On the night side, the pilot reported that the horizon was always visiblethrough the window. With full moon illumination, he stated that he couldaline the spacecraft in yaw nearly as well as on the daylight side. Thepilot was not able to use the eye patch to dark-adaptj he was able to seelittle, if any, of the ground or clouds before moonrise. The pilot reportedthat the periscope was not very useful, in general, for determining drifton the nightside. Even with a full moon, the clouds were too dim in theperiscope to pick up a specific point and follow it for determination ofyaw heading.

135



• •• •• • :.:

The pilot reported that he could use star drift as a reference at yawangles close to 90% but that within 10° of zero yaw, it was quite difficult.The pilot was also able to use constellations as a heading reference . Hereported that the number of stars he could see were approximately the same

as those normally seen from the ground on a dark night . He had no troublerecognizing constellations and therefore could use the stars to determine

heading by referring to his Star Navigation Charts.

Communication Activities

Table XVIII summarizes the astronaut ' s inflight communications. Thepilot reported that he believed more time was devoted to formal operationalreports during the MA- 6 mission than would be desirable for future flights.Reduction of the high proportion of communications involved in making radiocontact, reporting switch positions, and relaying instrument readings wouldpermit more detailed reports by the pilot of his activities and observations .

Scientific Observations

The astronaut's scientific observations are grouped into the followingcategories: celestial, meteorological, and terrestrial. In addition, thepilot's activities using the hand- held camera and a summary of subjectmaterial photographed are presented. See the Astronaut's Flight Reportand appendix C of reference 4 for additional information regardingscientific observations by the pilot and appendix D of reference 4 for ananalysis of these observations.

Celestial observations.- Numerous small particles, estimated by the

pilot to be from TT to ft inch in size, were observed during each sunrise periodas moving rearward past the spacecraft at a relative velocity estimated at1.3 to 2.2 meters per second. Since some of the particles were seen to driftinto the spacecraft's shadow, the estimated sizes and velocities can be reliedupon. The small relative velocity of the particles with respect to thespacecraft indicates that they were in nearly the same orbit, which makes itcertain that the particles came from the spacecraft. They undoubtedlydrifted because of the difference in aerodynamic drag. The most plausiblematerial for the particles is frozen water, since other substances are eithernot present in sufficient quantities, or are limited in source, which shouldbe exhausted by a third orbital pass. Water from the spacecraft environmental

control system is slightly more probable than water which was ejected as aproduct of the reaction within the control system thrusters because of the lowobserved velocities.

A luminous band observed around the horizon may be the result of internalreflections of the moonlit earth between two inclined windows in the space-craft . This explanation has been strengthened by observations of the band inearth photographs taken from the spacecraft, by calculations using theblueprints, and by direct observations in the trainer and in other spacecraft .The tan-to-buff color is found in one of ten observed reflections . If thisband was not a reflection, the pilot may have seen the 6300 and 6464 angstromred layer which is known to exist at about the altitude reported.

136



The sun was observed to be highly flattened on some photographs ofsunset, and this phenomenon was visually confirmed by the pilot. Spectra of sixstars in Orion were obtained with a hand-held, objective-prism spectrograph.

Meteorological observations.- A program of observations for the astronaut,suggested by U.S. Weather Bureau scientists, was designed to provide informa-

tion for the development of improved optical sensing equipment for satelliteweather-observations systems. Three of the requested observations reportedby the astronaut are listed below.

1. Determine whether cloud heights can be evaluated from orbital

altitude. Report: The astronaut reported he could identifycloud types and determine their heights.

2. Determine whether clouds can be seen on the dark side of theearth. Report: The pilot reported that with a full moon hewas able to see clouds and some vertical development on the

dark side.

3 - Determine whether lightning can be seen on the dark side of

the earth. Report: The astronaut reported that he couldclearly see lightning in two storms in the Indian Ocean, whichwas passed over in nighttime.

The Weather Bureau also suggested that pictures, using infrared filmand a special set of filters, be taken of cloud cover in order to evaluatethe relative effectiveness of various wave-length intervals for cloudobservation. The required filter and film were aboard the MA-6 flightspacecraft. However, because of operational requirements of a higherpriority, the pilot was not able to accomplish this exercise.

Terrestrial observations.- It is important to learn what the effectivevisual horizon for the pilot in orbital space may be and at what distancehe can recognize landmarks for use in navigation and attitude control. Thepilot's observations for the MA-6 flight are summarized in the followingparagraphs.

The astronaut confirmed what earlier earth-sky pictures and Tirossatellite photographs have indicated, that a large percentage of the earth'ssurface is covered by clouds. Only four land areas, the western African

desert, western Australia, the western United States, and the eastern coast

of the United States, were relatively free of clouds during the W hours of

flight.

The pilot reported that he could see the following landmarks during thedaylight periods across the United States: the cities of El Paso, Tex.,New Orleans, La., Charleston, S. C., and Savannah, Qa.j the Salton Sea inCalifornia; the Mississippi Delta; and Cape Canaveral, Fla. He also reporteda V-shaped figure in the water in the Atlantic, which he interpreted to bethe wake from a ship. This is probably the smallest object reported by the

137



•*• "• ! **; ; ••. **. • • ••• •• •••* • • • • • • • 7 - *.t*

pilot . If it was the wake of a destroyer, it would be approximately 120 feetin width and perhaps JOO yards long.

The only nonluminous feature seen at night, other than clouds, was afaint indication of the western coastline of Australia. Two types of featureswere reported, lightning produced by two storms in the Indian Ocean and the

lights of the city of Perth, Australia. The Indian Ocean ship flare andthe flares ignited at Woomera, Australia, were not seen, undoubtedly because

of cloud cover.

Color photographs .- The pilot was able to take a total of 70 photographs,38 of which were on one roll and the balance on a second roll. Table XIXcontains a complete listing of the subjects photographed and their approximatetimes of exposure . The pilot attempted to take pictures of the luminousparticles seen during sunrise using the color film. While there are severalphotographs containing specks, it cannot be definitely determined whetherthese specks are actually the particles observed or imperfections on thefilm and the window.

Sensation and Orientation During Weightlessness

General sensations.- The pilot reported that weightlessness was notunpleasant, caused no problem, and in several respects was easier or moreenjoyable than the Ig condition. For example, there were no pressure pointsfrom the seat, and certain tasks, such as using the camera and other personalequipment, were easier. This equipment could be left unsupported while anothertask was performed.

Of operational significance was his report that, under weightlessness,

the head assumes a new position because of the helmet tiedown straps. Thissuggests that the visual angles through the window and the periscope wouldbe slightly different than if they were measured from the fixed couchposition.

The pilot reported no problem in reaching for and activating controls.There was no tendency to overreach. This was expected, since experiencehas shown that the eyes will quickly correct for muscular reaction inaccu-racies .

Orientation.- The pilot reported that he experienced two illusions of

motion, but both of these illusions were associated with changing accelerationfields and not with weightlessness. The first occurred at sustainer enginecut-off (SECO), when he had a very slight feeling of pitching forward headover heels . The second illusion of motion occurred during retrograde, atwhich time he reported that he felt like he was going back towards Hawaii.He stated that following retrofire, when he was able to look out the windowand see the terrain moving away from him, this illusion disappeared.

No illusions of position occurred during weightlessness. The astronautdid not feel at any time that he was standing still and the earth was moving,or that he was in any position other than the true spacecraft positionduring the flight.

138



The pilot reported that he could sense angular accelerations that werepresent for a period of time sufficient to produce rates above 5

e/sec- This

is approximately in agreement with the results of tests on the pilot, conductedat Pensacola, Fla. . on the Human Disorientation Device (a two-dimensional ro-tational apparatus). There is no indication of greater sensitivity to ro-tational forces under zero g than Igj however, the observations are too limited

to warrant firm conclusion.

A final area of interest was his judgment of the vertical and thehorizontal during weightlessness. Normally, this perception is stronglyaffected by the otolith organs of the inner ear. Variations in the abilityto determine the horizontal have been demonstrated when the individual isasked to adjust a visible line in a completely dark room while lying on hisside. Errors in this adjustment usually occur after approximately 2 minutesin darkness. Provisions for conducting this test have been built into theminiaturized photometer carried aboard the spacecraft. The pilot's adjustmentto the horizontal was accurate. However, interpretation of the results is

difficult, since, because of the control system malfunction, the pilot washurried and made the line adjustment very shortly after looking into thedevice.

Personal Equipment

The equipment that the pilot used and his relevant comments arepresented in this section.

Daylight color camera.- A 35rom camera with a photoelectric cell toadjust the aperture automatically was provided. (See fig. 52(a).) It hasa 50mm, f-2.8 lens and has been provided with controls permitting rapidone-hand operation. The only comment regarding camera operation concernedchanging film. The pilot released the casette, and it floated behind theinstrument panel when he reached for it. The results of the photographyusing this camera are in the section labeled Color Photographs.

Ultra-violet spectogrsph.- A second 35™ camera with a special lenssystem adapted for ultra-violet spectral photography was included. (See

fig. 52(b).) The results of film experiment are inconclusive, and, therefore,of little value.

Photometer.- The photometer is a miniature device used by the pilot to

determine visually the intensity of celestia.1 light sources, to view thesun at sunset and to evaluate his capability to orient to the horizontal.(See fig. 52(c).)

Airglow filter.- All light except the 5577 angstrom line was filteredout by the airglow filter, (See fig. 52(d).) The only attempt to use thisfilter produced no results, primarily because of a low-level night adaptationat the time.

Night adaptation eye patch.- The form-fitted mold (see fig. 52(e)) whichis attached with tape prior to sunset worked well prior to lift-off, but it

139



I • •• •

failed to remain positioned during the flight. The combined effects of dust,humidity, and perspiration of the pilot reduced the effectiveness of the

adhesive.

Flight-plan cards.- Three cards, similar in style to figure 46, wereprovided to aid in maintaining schedules and to serve as a reminder of

upcoming events.

Food tube.- Two tubes were provided, one containing beef and vegetablesand the other applesauce. (See fig. 52(f).) The applesauce was consumedwithout difficulty, but the pilot did not have an opportunity to open theother tube.

Food tablets.- A food tablet dispenser was provided containing onexylose tablet and several malt tablets. (See fig. 52(g).)

140



• •• _»»

•• ••••••. • • •• • • ••• • • •

TABLE XV. - TIME EXPENDED IK ASTRONAUT PREIAUNCH ACTIVITIES

Date

January 1962

15

IT

19

20

23

27

29

Activity

Flight acceptance

composite test

launch simulation

Launch simulation

Simulated flight

Simulated flight

Countdown

Simulated flight

Total

Duration, hr:min(ai

7: 5

5:10

lj-:15

1:30

2:00

:00

1:15

25:55

Time does not include allowance for preparation, monitoring,debriefing, trouble shooting, or conferences.



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TABLE XVH. - CONTROL MODE USD ATTITUDE MANEUVERS DURING MA-6 MISSION

Control mode

(a)Elapsed time from launch,

hr:mln:sec

Attitude maneuvers

ASCS

MP, RSCS,

auxiliary damping,ra t

ASCS

EEW

ASCS

A S C S (pitch and roll),MP (yaw)

A S C S

FEW or MP

A S C S

nr

A S C S


A S C S (pitch and roll),FEW or MP (yaw)


ASCSA S C S and MP

MP

00:05:00 to 00:07:58

00:07:58 to 00:10:46

00:10:46 to 00:25:00

00:25:00 to 00:28:28

00:28:28 to 01:30:00

01:50:00 to 01:31:00

01:31:00 to 01:

01:31:17 to 01:

01:32:00 to 01:33:58

U31:17"\L:32:00 J

01:33:58 to 02:00:20

02:00:20 to 02:01:34^

02:01:34 to 02:01:42 )

02:01:42 to 02:02:08

02:02:08 to 92:02:23

02:02:23 to 02:08:29 102:08:29 to 02:13:28 \

02:13:28 to 03:31:00

Turnaround.

Control systems check. Cheeked MP, rate

c o m m a nd , auxiliary damping, and FEWfor

proper operation.

Held orbit attitude for radar tracking.

60' right-yawmaneuver to check periscope

as a yaw reference.

Left-yaw low-thrust failure commencing

at 01:29:30.

Started to drift out of orbit mode in yew;

brought it back by using MP.Eechecked left-yaw lowthrust; still faulty;

went to FEW or MP.

Drifted 35° in right yaw;went into orientation

mode; went back to FEW.

Day horizon check at 01:52:00; 180* rightyaw

maneuver from 01:53:27 to 01:59:10.

Rechecked A S C S operation; left-yaw low thruster

still faulty; checked operation of lowthrustersin yaw FEW and MP.

Right-yav lowthrugtera not operating; Intoorientation mode due to yawdrifting out to

left at 02:02:4?; switched to MP as aback-

up to A S C S

Orientation mode at 02:13:02; went to MP to

conserve fuel; held attitudes on Orion from

02:27:10 to 02:29:10; caged gyros byusing

night window reference at 02:40:18; uncaged

at approximately -10*in pitch and -20* in

roll;pitched down 60* to observe El Paso,Texas at 03:05:15; yawed 180* right from

03:14:00 to 03:29:00; caged and uncaged

gyros at 03:29:22, dayside; uncaged at

approximately -13* in pitch and 0* in roll

using window reference.

TCey:ASCS - Automatic stabilization and control systemRSCS - Rate stabilization control systemFEW - Fly-by-^fireMP - Manual proportional

143



• *• ••

TABLE XVH.- CONTROL MODE AND ATTITUDE MAKEOVERS DURING MA-6 MISSION - Concluded

Control mode

(a)

El a pse d time from launch,hr:mln:sec

Attitude maneuvers

ASCS

ASCS (pitch and roll),

MP (yaw)

ASCS and MP

FEW

ASCS

ASCS and MP

ASCS (pitch and roll),

MP (yaw)

MP

and MP

MP and ASCS or

auxiliary damping

03:J1:00 to 03:37:38

03:37:38to 03:lH:05

03:41:05 to 03:52:09

03:52:09 to 04:25:10

04:25:10 to 04:28:05

04:28:05 to 04:36:35

04:36:35 to 04:37:22

04:37:22 to 04:41:58

04:41:58 to 04:47:44

04:47:44 to 04:49:17

(drogue)

Rechecked ASCS operation.

Yawed left 35* to observe sunset.

ASCS with MP 'backup; slipped into orientation

mode at 03:51:20

Ho low thruBt in right yaw; held attitude on

Orion from 03:54:10 to 03:58:30;yawed 180°

right to observe sunset from 04:11:03 to

04:15:30; caged gyros at 04:21:26; uncaged at

approximately -14*in pitch and 0* in roll.

Rechecked ASCS.

ASCS with MP backup during retrosequence end

retrofire.

In and out of orientation due to faulty right

yaw thruster; controlled yaw manually.

Pitched down approximately 100* for geographi-

cal observations at 04:37:40; pitched manually

to reentry at 04:39:48.

Inserted at left roll of 12*/sec at 04:44:42;

effective damping to 04:47:00; pitch oscil-

lations pegged at 04:47:05; yew oscillations

pegged at 04:47:20.

Auxiliary damping or ASCS damped oscillations

from 04:47:44 to 04:48:40; rate oscillations

pegged in pitch and yaw again at 04:48:40

until antenna fairing separation.

"Key:ASCS - A u to m a ti c stabilization and control systemRSCS - Rate stabilization control systemS E W - Fly-by-wireM P - Manual proportional

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Figure 47.- One 60° right yaw maneuver using the periscope and fly-by-wire.

First maneuver (01:53=30)Second maneuver (03:14:00)

Third maneuver (04:11:00)

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Figure 48.- Three l80° right-yaw maneuvers using the window reference and manual

proportional or fly-by-wire.150



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Figure 52.- Concluded.

D : 155



PILOT'S FLIGHT REPORT

The pilot's report should be concerned mainly with those items in the

mission objectives where man's observation capabilities provide information

not attained by other means. It is in this type of reporting that a manned

vehicle provides a great advantage over an •unmanned vehicle, which is often

deaf and blind to the new and the unexpected. My report, then, will stress

what I heard, saw, and felt during the orbital mission.

Preparation and Countdown

Preparation, transfer to the launch complex, and insertion into the space-

craft went as planned. The technicians and I had been through the entry pro-

cedure into the spacecraft many times.

During the countdown, short but minor delays were encountered when problems

arose. The support for the microphone in the helmet, an item that had been

moved and adjusted literally thousands of times, broke and had to be replaced.

While the spacecraft hatch was being secured, a bolt was broken and had to be

repaired. During this time I was busy going over my checklist and monitoring

the spacecraft instruments.

The initial unusual experience of the mission is that of being on top of

the Atlas launch vehicle after the gantry has been pulled back. Through the

periscope, much of Cape Canaveral can be seen. If you move back and forth in

the couch, you can feel the entire vehicle moving very slightly. When the

launch-vehicle engines are gimbaled, you can feel the vibration, and when thetank is filled with liquid oxygen, the spacecraft vibrates and shudders as the

metal skin flexes. The white plume of the vented lox (liquid oxygen) is visible

through the window and periscope.

Powered Flight

When the countdown had reached zero, I could feel the engines start. The

spacecraft shook, not violently but very solidly. There was no doubt when lift-

off occurred. When the launch-vehicle was released, there was an immediate

gentle surge that let me know I was on my way. The roll to the correct azimuth

was noticeable after lift-off. I had preset the small window mirror to watch

the ground. I glanced up after lift-off and could see the horizon rotating.

Some vibration had occurred immediately after lift-off, but it smoothed out

after about 10 to 15 seconds of flight. There was still a noticeable amount

of vibration that continued up to the time the spacecraft passed through the

maximum aerodynamic pressure q . The approach of maximum q. is signaled by

more intense vibrations. During this period, I was conscious of a dull muffled

roar from the engines. Beyond this period the vibration smoothed out noticeably.

However, the spacecraft never became completely free of vibration during powered

flight.

156



The acceleration buildup was noticeable but not bothersome. Booster

engine cut-off occurred at 00:02:09, and as the two outboard engines shut down

and were detached, the acceleration dropped but not as sharply as I had antic-

ipated. Instead, it decayed over approximately 1/2 second. There is a changein noise level and vibration when these engines are jettisoned. I saw a flashof smoke out the window and thought at first that the escape tower had jetti-

soned early, which I reported. However, this flash was apparently deflectedsmoke coming up around the spacecraft from the booster engines which had .lust

separated. The tower was jettisoned at 00:02:33*an<i I corrected my earlier

report. I was" ready to back up the automatic sequencing system if it did notperform correctly and counted down the seconds to the time for tower jettisoning.I was looking at the nozzles of the tower rockets when they were ignited. Alarge cloud of smoke came out but little flame. The tower accelerated rapidlyfrom the spacecraft in a straight line. I watched it to a distance of approxi-mately 1/2 mile. The spacecraft was programed to pitch down slowly just prior'to jettisoning the tower, and this maneuver provided my first real view of thehorizon and clouds.

After the tower was jettisoned, the spacecraft pitched slowly up again1and

I lost sight of the horizon. I remember making a comment at about this timethat the sky was very black. I could communicate well up to and during the

time of the maximum acceleration of 7.7g, which occurred at sustainer engine

cut-off (SECO).

Just before the end of powered flight, there was one experience I was notexpecting. At this time the fuel and lox tanks were nearly empty, and appar-

ently the launch vehicle becomes considerably more flexible under these cond-itions than when it is filled. • I had the -sensation of being out on the end of aspring-board and could feel oscillating motions as if the nose of the launch

vehicle were waving back and forth slightly-

Orbital Insertion

The noise level also increased as the vehicle approached SECO. When SECO

occurred at 00:05:01 and the acceleration dropped to zero, I had a veryslight sensation of tumbling forward. The astronauts have often had a similarsensation during training on the centrifuge, but the sensation was much lesspronounced during the flight.

There was no doubt when the explosion bolts holding the clamp ring between

the launch vehicle and the spacecraft were ignited. There was a loud reportand I immediately felt the force of the posigrade rockets which separate thespacecraft from the launch vehicle and provide the final insertion impulse.Prior to the flight I had imagined that the acceleration from these three small

rockets would be insignificant and that I might fail to sense them entirely,but there was no doubt when they ignited.

As the spacecraft came around to its normal aft viewing attitude, I could

see the launch-vehicle sustainer stage through the window. At that time, Iestimated that it was "a couple of hundred yards away. " After the flight an

analysis of the trajectory data showed this distance to be 600 feet. The

157



capability to estimate distances of this magnitude will be important in futuremissions in which the pilot will want to achieve rendezvous, since he will becounted on to perform the final closing maneuver.

t

I was able to keep the sustainer stage in sight for 6 or 7 minutes whileit traveled over the Atlantic. At the last time I reported seeing it, it was

approximately 2 miles behind and 1 mile below the spacecraft. It could be seeneasily as a bright silvery object against the black background of space.

Orbit

The control system turned the spacecraft around and oriented it into the

proper attitude. After my initial contact with Bermuda, I received the times

for firing the retrorockets and conducted a check of the controls. I hadpracticed it many times on the ground in the Mercury procedures trainer and thetest went just as it had in the trainer. This experience was the first time Ihad been in complete manual control of the spacecraft, and it was very reas-

suring to see not only the spacecraft react as expected, but also to see thatmy own ability to control was as we had hoped.

Following this controls check, I went back to automatic stabilization andcontrol system (ASCS) control and the spacecraft operated properly in this modethroughout the first orbital pass.

Thruster problem. - Because of a malfunction in a low-torque thruster atthe end ,pf the first orbital pass, it was necessary for me to control the space-craft manually for the last two passes. This requirement introduced no seriousproblems, and it actually provided me with an opportunity to demonstrate whata man can do in controlling a spacecraft. However, it limited the time thatcould be spent on many of the experiments and planned observations I had hopedto carry out during the flight.

Attitude reference.- A number of questions have been raised over theability of man to use the earth's horizon as a reference for controlling theattitude of the space vehicle.

Throughout this flight, no trouble in seeing the horizon was encountered.During the day the earth is bright and the background of space is dark. Thehorizon is vividly marked. At night, before the moon is up, the horizon canstill be seen against the background of stars. After the moon rises (during

this flight the moon was full), the earth is well enough lighted so that thehorizon can be clearly seen.

With the lighted horizon as a reference, the pitch and roll attitudes ofthe spacecraft can easily be controlled. Yaw, or heading reference, however,is not so good. I believe that there was a learning period during the flightregarding my ability to determine attitudes in yaw. Use of the view throughthe window and periscope for this purpose gradually improved.

To determine yaw attitude, advantage must be taken of the speed of thespacecraft over the earth which produces an apparent drift of.the ground below

158



the spacecraft. When the spacecraft is oriented in the plane of the orbit,the ground appears to move parallel to the spacecraft longitudinal axis.During the flight I developed a procedure which seemed to help me use thisterrain drift as a. yaw reference. I would pitch the small end of the space-craft down to about -60° from the normal attitude where a fairly good verticalview was available. In this attitude, clouds and land moving out from under me

had more apparent motion than when the spacecraft was in its normal orbit atti-tude pointing toward the horizon.

At night with the full moon illuminating the clouds below, I could stilldetermine yaw attitude through the window but not as rapidly as in the daytime.At night, I could also use the drift of the stars to determine\heading, although

this procedure took longer and was less accurate.

Throughout the flight, I preferred the window to the periscope as an atti-tude reference system. It seemed to take longer to adjust yaw by using theperiscope on the day side. At night, the cloud illumination by the moon is too

dim to be seen well through the periscope.

Three times during the flight I turned the spacecraft approximately l80° inyaw and faced forward in the direction of flight. I liked this attitude - seeingwhere I was going rather than where I had .been - much better. As a result ofthese maneuvers my instrument reference system gave me an inaccurate attitudeindication. It was easy to determine the proper attitude, however, from refer-

ence to the horizon through the window or the periscope. Maintaining orien-tation was no problem, but I believe that the pilot automatically relies muchmore completely on vision in space than he does in an airplane, where gravitycues are available. The success with which I was able to control the space-craft at all times was one of the most significant features of the flight.

Weightlessness.- The sensation of weightlessness was a pleasant experience.I reported that I felt fine as soon as the spacecraft separated from the launchvehicle, and throughout the flight this feeling continued to be the same.

Approximately every 30 minutes throughout the flight I went through-aseries of exercises to determine whether weightlessness was affecting me in anyway. To see if head movement in a zero-g -environment produced any symptoms ofnausea or vertigo, I tried first moving, then shaking my head from side to side,up and down, and tilting it from shoulder to shoulder^ in other words, movingmy head in roll, pitch, and yaw. I began slowly, but as the flight progressed,

I moved my head more rapidly and vigorously until, at the end of the flight, Iwas moving rapidly as my pressure suit would allow.

In another test, using only eye motions, I tracked a rapidly moving spotof light generated by my finger-tip flashlights. I had no problem in watchingthe spot and once again no sensations of dizziness or nausea. A small eyechart was included on the instrument panel with letters of varying size andwith a "spoked wheel" pattern to check both general vision and any tendencytoward astigmatism.. No change from normal was apparent.

159



•• ••• • ••• •> • • • • •• • •• • •• •

An "oculogyric test" was also made in which turning rates of the spacecraftwere correlated with sensations and eye movements. These results were againnormal. Preflight experience in this test had been conducted and a calibrationhad been made at the Naval School of Aviation Medicine, Pensacola, Florida, sothat I was thoroughly familiar with my reactions to these same movements at Ig.

To provide medical data on the cardiovascular system, I conducted period-

ically an exercise which consisted of -pulling on a bungee cord once a secondfor JO seconds. This exercise provided a known workload to compare with previ-ous similar tests made on the ground. The flight surgeons have reported theeffect that this had on my pulse and blood pressure. The effect that it had onme during the flight was the same effect that is had on the ground - it made metired.

Another experiment related.to the possible medical effects of weightless-ness was eating in orbit. On the relatively short flight of Friendship"J,eating was not a necessity, but rather an attempt to determine whether there

would be any problem in consuming and digesting food in a weightless state. Atno time did I have any difficulty eating. I believe that any type of food canbe eaten as long as it does not come apart easily or make crumbs.

Sitting in the spacecraft under zero g is more pleasant than under Ig onthe ground, since you are not subject to any pressure points. I adapted quiterapidly to weightlessness. I had no tendency to overreach and did not experi-

ence any other sign of lack of coordination, even on the first movements afterspacecraft separation. I found myself unconsciously taking advantage of theweightless condition, such as when I would leave a camera or some other objecthanging motionless while I attended to other matters. This procedure was not

done as a preplanned maneuver but as a spontaneous and natural thing when otheritems required my attention. I thought later about how I had done this asnaturally as if I were laying the camera on a table in a Ig field. It pointedlyillustrates how rapidly adaptable the -human is, even to something as foreignas weightlessness.

I had brought along a number of instruments, such as cameras, binoculars,and a photometer, with which to make observations from a spacecraft. All ofthese items were stowed in a special equipment storage kit by my right arm.Each piece of equipment had a J-foot piece of line attached to it. By the time

I had started using items of the equipment, these lines became tangled. Al-though these lines got in the way, it was still important to have some way of

securing the equipment, as I found out when I attempted to change film. Thesmall canisters of film were not tied to the equipment kit by lines. I leftone hanging motionless while working with the camera. When I reached for it,I accidentally hit it and 'it floated out of sight behind the instrument panel.

Color, light, and visibility.- As I looked back at the earth from space,colors and light intensities were much the same as I had observed when flyingat high altitude in an airplane. When looking toward the horizon, however, theview is completely different, for then the blackness of space contrasts vividlywith the brightness of the earth. The horizon itself is a brilliant blue and

white.

160



It was surprising how much of the earth's surface was. covered by clouds.The clouds can be seen very clearly on the daylight side. The different typesof clouds - vertical developments, stratus clouds, and cumulus clouds - arereadily distinguishable. There is little problem in identifying them or inseeing the weather patterns. One can estimate the relative heights of thecloud layers from a knowledge of the cloud types or from the shadows that highclouds cast on those below them. The U.S. Weather Bureau was interested in im-proving the optical equipment in their Tiros and Nimbus satellites and wantedto know if I could determine the altitude of cloud layers with better opticalresolution. From my flight, it seemed quite possible to determine cloud heightsfrom orbital altitudes.

Only a few land areas were visible during the flight .because of the cloudcover. Clouds were over much of the Atlantic, but the western (Sahara Desert)part of Africa was clear. In this desert region I could plainly see dust storms.By the time I got to the east coast of Africa where I might have been able tosee towns, the land was covered by clouds. The Indian Ocean was the same.

Western Austrailia was clear, but the.eastern half was overcast. Most ofthe area across Mexico and nearly to New Orleans was covered with high cirrusclouds. As I came across the United States I could see New Orleans, Charleston,and Savannah very clearly. I could also see rivers and lakes. . I think the bestview I had of any land area during the flight was the clear desert regionaround El Paso on the second pass across the United States. I could see thecolors of the desert and the irrigated area north of El Paso. As I passed offthe east coast of the United States I could see across Florida and far backalong the Gulf Coast.

Over the Atlantic I saw what I assume was the Gulf Stream. The differentcolors of the water were clearly visible.

I also observed what was probably the wake of a ship. As 1 was passingover the recovery area at the end of the second orbit, I looked down at thewater and saw a little "V. " I checked the map. I was over recovery area G atthe time, so I think it was probably the wake from a recovery ship. When I-looked again the little "V" was under a cloud. The change in light reflectionscaused by the wake of a ship are sometimes visible for long distances from anairplane and will linger for miles behind a ship.

I believe,, however, that most people have an erroneous conception that

from orbital altitude, little detail can be seen. On the ground, it is commonto see a mountain range, in clear desert air 100 or so miles away very clearly,and all.that vision is through.dense atmosphere. From orbital altitudes, at-mospheric light attenuation is only through approximately 100,000 feet ofgradually less dense atmosphere so it is even more clear.

Obviously, on the night side of the earth, much less was visible. Thisfact may have been caused not only by the reduced light, but also by the factthat I. was never fully dark adapted. In the bright light of the full moon, theclouds are visible, and I could see vertical development at night. Most .of thecloudy areas, however, appeared to be stratoform.

161



•••••

The lights of the city of Perth, in western Australia, were on and I couldsee them well. The view was similar to that seen when flying at high altitude

at night over a small town. South of Perth there was a small group of lights,but they were much brighter in intensity. Inland there was a series of four orfive towns lying in a line running from east to west. Knowing that Perth was

on the coast, I was just barely able to see the coastline of Australia. Cloudscovered the area of eastern Australia around Woomera, and I saw nothing butclouds from there across the Pacific until I was east of Hawaii. There ap-peared to be almost solid cloud cover all the way.

Just off the east coast of Africa, there were two large storm areas.Weather Bureau scientists had wondered whether lightning could be seen on thenight side, and it certainly can. A large storm was visible just north of my

orbital track over the Indian Ocean and a smaller one to the south. Lightningcould be seen flashing back and forth between the clouds but most prominent.were lightening flashes within thunderheads illuminating them like light bulbs.

Some of the most spectacular sights during the flight were sunsets. Thesunsets always occurred slightly to my left, and I turned the spacecraft to geta better view. The sunlight coming in the window was very brilliant, with anintense clear white light that reminded me of the arc lights while the space-craft was on the launching pad.

I watched the first sunset through the photometer, which had a polarizingfilter on the front so that the intensity of the sun could be reduced to acomfortable level for viewing. Later I found that by squinting, I could lookdirectly at the sun with no ill effects, just as I can from the surface of theearth. This accomplished little of value but does give an idea of intensity.

The sun is perfectly round as it approaches the horizon. It remainscircular until it reaches the horizon, and retains its symmetry until justthe last sliver is visible. The horizon on each side of the sun is extremelybright, and when - t h e sun has gone down to the level of this bright band of thehorizon, it seems to spread out ^5° "to 60° to each side of the point where it issetting. With the camera I caught the flattening of the sun just before it set.

As the sun moves toward the horizon, a black shadow of darkness movesacross the earth until the whole surface, except for the bright band at thehorizon, is dark. This band is extremely bright just as the sun sets, but astime passes the bottom layer becomes a bright orange and fades into reds, thenon-into the darker colors, and finally off into the blues and blacks. One thingthat surprised me was the distance that the light extends on the horizon oneach side of the point of the sunset. The eye can see a little 'more of thesunset color band than a camera captures. One point of interest was the lengthof time during which the orbital twilight persisted. Light was visible alongthe horizon for k to 5 minutes after the sunset, which is a long time when youconsider that sunset occurred 18 times faster than normal.

The period immediately following sunset was of special interest to theastronomers. Because of atmospheric light scattering, it is not possible tostudy the region close to the sun except at the time of a solar eclipse. It

162



had been hoped that from above the atmosphere the area close to the sun couldbe observed. However, this observation would require a period of dark adapta-tion prior to sunset. An eye patch, which was to be held in place by specialtape, had been adapted for this purpose. The patch was expected to permit oneeye to be night adapted prior to sunset. Unfortunately, the tape proved unsat-

isfactory and I could not effectively use the eyepatch. Observations of thesun's corona and zodiacal light must await future flights when the pilot mayhave an opportunity to get more fully dark adapted prior to sunset.

Space particles.- The biggest surprise of the flight occurred at dawn.Coming out of the night on the first pass, at the first glint of sunlight onthe spacecraft, I was looking inside the spacecraft checking instruments forperhaps 15 to 20 seconds. When I glanced back through the window my initialreaction was that the spacecraft had tumbled and that I could see nothing butstars through the window. I realized, however, that I was still in the normalattitude and that the spacecraft was surrounded by luminous particles.

These particles were a light yellowish green color. It was as if thespacecraft were moving through a field of fireflies. They were about thebrightness of a first magnitude star and appeared to vary in size from a pin-head up to possibly 3/8 inch. They were about 8 to 10 feet apart and evenlydistributed throughout the space immediately around the spacecraft. Occa-sionally, one or two of them would move slowly up around the spacecraft andacross the window, drifting very, very slowly, and would then gradually moveoff, back in the direction in which I was looking. I observed these luminousobjects for approximately k- minutes during each sunrise period.

During the third sunrise, I turned the spacecraft around and faced forward

to see if I could determine where the particles were coming from. Facing inthis forward direction I could see only about 10 percent as many particles as Ihad seen when my back was to the sun. Still, they seemed to be coming towardsme from some distance so that they appeared not to be coming from the space-craft. Just what these particles are is still subject to debate and awaitsfurther clarification.

Other planned observations. - A number of other observations and measure-ments during orbit had to be canceled because of the control .system problems.The scientific equipment carried aloft was not highly sophisticated. It wasbelieved, however, that it would show the feasibility of making. more compre-hensive measurements on later missions.

• Some of these areas of investigation that we planned but did not have an

opportunity to check are as follows:

(a) Weather Bureau observations:(1) Photographs of weather areas and cloud formations to match

against map forecasts and Tiros photographs.

(2) Filter'mosaic photographs of major weather centers.

(3) .Observation of .green airglow from air and weather centers• .in 5577-angstrom band.with an airglow filter.

163



n " i i(k) Albedo intensities .- -measure reflected light intensities on

both day and night side

(b) Astronomical observations:(1) Light polarization from the area of sun

(2) Comets close to sun(3) Zodiacal light

(4) Sunlight intensity

(5) Lunar clouds

(6) Gegenschein

(7) Starlight intensity measurements

(c) Test for otolith balance disturbance and autokinesis phenomena

(d) Vision tests:(1) Night vision adaptation

(2) Photometer and eye measurements

(e) Drinking

Reentry

After having turned around on the last orbital pass to see the spaceparticles, I maneuvered into the correct attitude for igniting the retrorockets

and stowed the equipment in the equipment storage kit.

During the last dawn, my attitude indicators were still slightly in error.However, before it was time to ignite the retrorockets, the horizon-scannerslaving mechanism had brought the gyros back to orbit attitude. I crosscheckedrepeatedly between the instruments, periscope presentation, and the attitudethrough the window.

Although there were variations in "cne instrument presentations during theflight, there was never any difficulty in determining my true attitude by ref-erence to the window or periscope. I received a countdown from the ground andthe retrorockets were ignited on schedule just off the California coast.

I could hear the report of each rocket and could feel the surge as therockets slowed the spacecraft. Coming out of zero-g condition, the force ofthe retrorockets produced the sensation that I was accelerating back towardHawaii. This sensation, of course, was an illusion.

Following retrofire, the decision was made to have me reenter with the

retropackage still attached because of the uncertainty as to whether the heat-shield had been released. This decision required me to perform manually anumber of.the -operations which are normally conducted automatically during thereentry. I brought the spacecraft to the proper attitude for reentry using

164



manual control. The periscope was retracted by pumping the manual reactionlever.

As deceleration began to increase I could hear a definite hissing noisethat sounded like small particles brushing against the spacecraft.

Because of the ionization of the air around the spacecraft, communica-tions were lost. This loss had occurred on earlier missions and was expected-

As the heat pulse started, there was.a noise and a bump on the spacecraft. Isaw one of the straps that hold the retropackage swing in front of the window.

The heat pulse increased until I could see a glowing .orange color throughthe window. Flaming pieces were breaking off and flying past the spacecraftwindow. At the time, these observations were of some concern to me becauseI was not sure what they were. I had assumed that the retropackage had beenjettisoned when I saw the strap in front of the window. I thought theseflaming pieces might be pa'rts of the heat shield breaking off. I know now,

of course, that the pieces were from the retropackage.

There was no doubt when the heat pulse occurred during reentry but it

takes time for the heat to soak into the spacecraft and heat the air. I didnot feel particularly hot until I had descended to about 75,000 to 80,000 feet.

From there on down I was uncomfortably warm, and by the time the main parachute

was out I was perspiring profusely.

The reentry deceleration of 7-Tg wa-

s as expected and was similar to thatexperienced in centrifuge tests. There had been some question as to whether

our ability to tolerate acceleration might be worse because of the hours

of weightlessness, but I could note no difference between my feelings ofdeceleration of this flight and my training sessions in the centrifuge.

After peak deceleration, the amplitude of the spacecraft oscillationsbegan to build. I kept them under control in the manual proportional andfly-by-wire modes until I ran out of manual fuel. After that point, I wasunknowingly left with only the fly-by-wire system and the oscillations in-creased; so I switched to auxiliary damping, which controlled the spacecraftuntil the automatic fuel was also expended. I was reaching for the switchto deploy the drogue parachute early in order to reduce these reentry oscil-lations, when it was deployed automatically. The drogue parachute stabilized

the spacecraft rapidly.

At 10,800 feet the main parachute was deployed. I could see it streamout behind me momentarily, fill partially, and then as the reefing line cutterswere actuated iVfilled completely. The opening of the parachute caused ajolt, but perhaps less than I expected.

165



•• ••• • ••• • •• •••• •• •• • ••••• ••• •* • • . . ;•• •• ••• • »•»

Landing and Recovery

The landing deceleration vas sharper than I"had expected. Prior tolanding I had disconnected all the extra leads to my pressure suit and wasready for rapid egress, T a u t there was no need for t h i s . - I had a messagethat the destroyer U.S.S..Noa would pick me up within 20 minutes. I layquietly in the spacecraft trying to keep as cool' as possible. The tempera-ture inside the spacecraft did not seem to diminish. This temperature level,

combined with the high humidity of the air being drawn into the -spacecraft,

kept me uncomfortably warm and perspiring heavily. Once the U.S.S. Noa'wasalongside the spacecraft, there was little delay in starting the hoisting

operation. The spacecraft-was pulled part way out of the water to let thewater drain from the landing bag.

During the spacecraft pickup, I received one good bump.- It was probablythe most solid jolt of the whole trip as -the spacecraft swung against theside of the ship. Shortly afterwards the spacecraft was on the deck.

I had initially planned egress out through the top, but by this time Ihad been perspiring heavily for nearly ^5 minutes.' I decided to come outthe side hatch instead.

- Concluding Remarks

The first orbital flight of a manned Mercury spacecraft has proved thatman can adapt very rapidly to the space /environment. My senses and capabil-

ities appeared unchanged in space, at least for the ^^-hour duration of

thisflight, and

weightlessnesswas no

problem.

Of major significance is. the fact that much more dependence can beplaced on the man as a reliably operating portion of the man-spacecraftsystem. In many areas, his safe return can be made dependent on his ownintelligent actions. Even where automatic systems are still necessary,mission reliability is~ tremendously increased by having the man as a backup.

Man's adaptability is most evident in his powers of observation. He can

accomplish many-more and varied experiments during each mission than can beobtained from an unmanned vehicle. When the unexpected arises, as happenedwith the luminous particles -and horizon-band observations on the flight, he

can make observations that will permit a more rapid evaluation of thesephenomena on future flights. Indeed, on an unmanned flight there likelywould have been no such observations.

On the ground some things might be done differently. As an example, itwould be more advisable in the event of suspected malfunctions, such as theheat-shieldlanding-bag difficulties, which require extensive discussion tokeep the pilot updated on each bit of information, rather than waiting to

166



. •

send a final clearcut recommendation from the ground. This procedure keeps thepilot fully informed at all times, which is important if there should happento be any communication difficulty and it became necessary for him to make alldecisions solely from onboard information.

Most important, however, the future will not always find us as powerlimited as we are now. We will progress to the point where missions will notbe totally preplanned. Therefore, a greater number of alternatives will be

available for the pilot to choose from during a flight, and consequently man'sintelligence and decision-making capability will become mandatory.

167



••••• • •!

• • • •

• • •• • «

• • • •

••••• •

LAUNCH-VEHICLE PERFORMANCE

All launch-vehicle systems performed satisfactorily. The followingitems are noted for information.

.Abort Sensing and Implementation System (ASIS)

The ASIS performed satisfactorily. Rone of the abort parameters werenear abort threshold. As expected in normal sequence, an ASIS abort signalwas generated following sustainer engine cutoff (SECO).

Engine Cutoff

SECO and auxiliary sustainer cutoff (ASCO) signals were transmitted,and at least one was received and properly acted upon by the launch vehicle.

Instrumentation does not permit determination of whether or not both signalswere acted upon by the launch vehicle.

Orbit Lifetime

Computed data based upon probable thrust having been imparted to thelaunch-vehicle tankage by the spacecraft posigrade rockets (-k ft/sec)

indicated at least 10 orbital passes to be expected from the launch vehicle.The final stage tankage, however, was later found to have reentered some6 passes following launch. Tracking during the third orbital pass indicateda perigee of about 95 nautical miles, an apogee of about 131 nautical miles,

and a period of approximately 8j minutes. No useful tracking data were

obtained after the fourth orbital pass.

Guidance

The performance of the Atlas launch-vehicle radio guidance system wasexcellent in this flight. The guidance system locked-on the vehicle inboth track and rate at 00:00:68, approximately as planned, and lost lockat 00:05:32 (2.0 seconds after SECO). Launch-vehicle oscillations resultingfrom guidance-system pitch-steering commands are shown in figure 53 andwere within acceptable limits. These rates were measured by the spacecraftrate gyros, and the data from the MA-5 mission are shown for comparison.

In figures 5^ to 56, the velocity and flight-path angle are shown inthe region of sustainer cutoff.. Guidance system data are shown in figure 54,and the Azusa data used in the Range Safety Impact Predictor Computer(IP 7090) are shown in figure 55 to illustrate the noise level during thetime of the go-no-go computations. Both the launch-vehicle radar and theIP 7090 data are considered excellent, except for two Azusa points im-mediately after SECO. The cause of these Azusa points scattering has beentraced to a recorder in the line. Except for these two points, the datascatter shown in the figures is considered normal and is smoothed or

168



averaged for the go-no-go decision.

The system gave a cutoff condition which was about 7 ft/sec low invelocity and about 0-50° low in flight-path angle. These values are withinthe expected accuracy range for the system. In figure 56, these data areshown as flight-path angle plotted against velocity. This is the type of

display used by the Flight Dynamics Officer in the Mercury Control Centerfor the orbital gono-go decision. Both the launch-vehicle guidance systemand IP 7090 data indicated a go condition.

. Aerodynamic Loads

The angle of attack times dynamic pressure aq for 'the flight is shownin figure.57 and is based on the measured wind profile at launch.

TRAJECTORY AND MISSION EVENTS . ' , .

. , Sequence o f Flight Events .

Comparisons of planned .and. actual major event times and pertinenttrajectory information are presented in table XX for the MA-6launch, -• .orbital, and reentry phases. These data generally reflect launch-vehicleguidance accuracy and the performance of the spacecraft sequential systemin effecting these flight events. -

The parameters shown for the "planned" launch trajectory in table XXwere computed by using the 1959 ARDC model atmosphere to maintain consistencywith preflight trajectory computations calculated before the flight. The

density of the atmosphere at Cape Canaveral is approximately 10 percenthigher than that of ARDC model atmosphere in the r.egion of maximum dy-namic pressure (an altitude of about 37>000 feet). As a result, themaximum dynamic pressure expected would be about 10 percent higher than thatshown as "planned." For this flight, the maximum dynamic pressure experiencedwas about 12 percent higher than that shown as "planned."

Flight Trajectory . •

The entire MA-6flight trajectory, from lift-off through landing, wasdetermined and .is compared with the nominal case. Figure 58 presents the

profile of altitude plotted against longitude,for the entire flight, anddetailed data are given below.

Launch phase.- Launch trajectory data, shown.in figure 59^ are basedon the real-time output of the Range Safety Impact Predictor Computer (whichused Azusa MK II and Cape Canaveral FPS-16 radar data) and the launch-vehicleguidance computer. The data from these tracking facilities were used duringthe time periods listed on following page. ,

169



Facility

Cape Canaveral FPS-16

Azusa MK II

Launch-vehicle guidance system

Time, min:sec

0 to 00:45

00:45 to 01:08

01:08 to 05:02

Orbital phase.- The orbital portion of the trajectory, shown infigure 60, was derived first by starting with the spacecraft position andvelocity vector obtained during the second orbital pass near Woomera,Australia, as determined by the Goddard computer using Mercury Networktracking data. Integration backward along the flight trajectory toorbital insertion and forward to the start of retrofire at the end of thethird pass yielded the calculated orbital trajectory. These integrated

values were in excellent agreement with the measured values of the guidancesystem at orbital insertion. They were also in accord with the positionand velocity vectors determined by the Goddard computer for passes near theCanary Islands (first pass), Bermuda (second and third passes), and Muchea,Australia (during the third pass), thus establishing the validity of theintegrated orbital portion of the flight trajectory.

Reentry phase.- The reentry portion of the trajectory, shown infigure 6l, was obtained by starting with the spacecraft position and velocityvector near Corpus Christ!, Texas, as determined by the Goddard computer.Integration backward along the flight to the end of retrofire and forwardto landing.yielded the reentry trajectory. This exercise assumed that the

retropackage had not been jettisoned and that the drogue parachute wasdeployed at 04:49:17 (given by telemetry) at an altitude of approximately27,000 feet, instead of the planned altitude of 21,000 feet. The spacecraftdecelerations from the integrated reentry trajectory agree within readingaccuracy with the decelerations measured by the onboard accelerometer. Inaddition, the times of drogue and main parachute deployment from the integratedreentry trajectory and those taken from spacecraft onboard measurements agreewithin 1 second. This agreement with measurements recorded in the space-craft serves to verify the validity of the integrated reentry portion of thetrajectory. The integrated values at the end of retrofire were adjustedby adding the effects of a nominal retrorocket total impulse of 38>880 Ib-sec

at nominal spacecraft retrofire attitudes of -34° in pitch and with 0° inroll andyaw.

The results, when compared with the orbital integrated values at thestart of retrofire, show that the velocity was low by about 7 ft/sec. Thisindicates that the velocity increment actually imparted to the spacecraft atretrofire was in excess of its nominal value by an equal amount. .Thisexcessive velocity can result from two factors: the weight of the space-craft having been below its nominal magnitude during retrograde and theretrorocket thrust having been above normal. The fact that the spacecraftlanded approximately 40 nautical miles short of the expected landing point

170



can primarily be attributed to this abnormally high velocity increment,with an additional but lesser effect resulting from the retrograde attitudebeing slightly above normal. An error of 1 ft/sec in the velocity incrementat retrofire will give a corresponding error in landing range of 5-2 nauticalmiles from the nominal landing point, and an error of 1.0° in pitch attitudeduring retrofire will give an error in landing range of 10.0 nautical miles

from nominal. The reentry trajectory and the landing point were only slightlyaffected by the retention of the retropackage. The integrated landing pointwas about 4 nautical miles.short of the spacecraft pickup point.

The aerodynamic parameters for the planned and integrated reentrytrajectories were computed by using the HASA Manned Spacecraft Center modelatmosphere. This is based on Discoverer Satellite program data foraltitudes above 50 nautical miles, the 1959 ARDC model atmosphere foraltitudes between 25 and 50 nautical miles, and the Patrick Air Force Baseatmosphere for altitudes below 25 nautical miles.

In the trajectory figures (figs. 59 to 6l), the above integrated valuesare labeled "actual." All planned values presented were precisely calcula-ted before the flight for the known mission parameters to arrive at theexpected trajectory data.

A comparison of the planned and actual trajectory parameters is givenin table XXI. The difference between these values primarily resulted fromthe actual cutoff velocity and flight-path angle at insertion being slightlylower than planned.

171



• ••• • «. „, .•• • . . • . .::•::• • • •. . ... • ««.

TABLE XX. SEQUENCE OF EVENTS

Event£ L

Planned time,hr:min: sec

Actual time,hr:min:sec

Difference,sec

Launch phase

Booster engine cutoff (BECO)

Tower release

Escape rocket Ignition

Sustainer-engine cutoffdiscrete (SECO)

Tail-off complete

Spacecraft separation

00:02:11.4

00:02:34.2

00:02:34.2

00:05:03.8

00:05: 03. 8

00: 02: 09. 6

00:02:33.3

00: 02: 33. 4

00: 05: 02

00:05:02

00:05:03.6

1.8

-0.9

-0.8

-1.8

-0.2

Orbital phase

Retrograde initiation

Retrorocket no. 1 (left)

Retrorocket no. 2 (bottom)

Retrorocket no. 3 (right)

Retropackage jettison

04:32:58

04:32:58

04:33:03

04:33:08

04:33=58

04:33:08

04:33:08

04:33:13

04:33.18

Reentry phase

0. 05g relay actuation

Drogue parachute deployment

Main parachute deployment

Main parachute jettison

04:43:53

04:50:00

04:50:36

04: 55: 22

b04:43:31

04:49:17.2

04: 50: 11

04: 55: 23

J.0.0

10.0

10.0

10.0

-22.0C(0)

-42.8C(-1.0)

-25.0C(-1.0)

1.0C(-27.0)

aPreflight calculated, based on nominal launch-vehicle performance.

The 0.05g relay was actuated manually by the astronaut when he was in a"small g field."

Q

Numbers in parentheses show the time difference between the actual eventbased on insertion parameters and the postflight-calculated reentry event.

172



• •• *

TABLE XXI.- COMPARISON OF PLANNED AND ACTUAL TRAJECTORY PARAMETERS

Condition and quantity Planned |Actual |Difference

Cutoff conditions (including tail-off)

Range time, sec 303-8 302.0

Range time, mln:sec 05:OJ.8 05:02

Geodetic latitude, deg North 30.4273 30.1*533

Longitude, deg West 72-5268 72-5865

Altitude, ft 528,428 528,381

Altitude, nautical miles 86.97 86.96

Range, nautical miles 436.4 1*33-7

Space-fixed -velocity, ft/sec 25,715 25,708

Space-fixed flight-path angle, deg 0 -0.01*68

Space-fixed heading angle, deg East of North 77-4756 77.1*826

Post-posigrade firing conditions

Range time, sec 305-3 306.8

Range time, min: sec 05:05.8 05:06.8

Geodetic latitude, deg North 30.1*572 JO. 5128

Longitude, deg West 72.3797 72.2923

Altitude, ft 528,1*60 528,361

Altitude, nautical miles 86.98 86.96

Range, nautical miles 1*1*4.2 1*49.4

Space-fixed velocity, ft/sec 25,737 25,730

Space-fixed flight-path angle, deg -0.0030 -0.0517

Space-fixed heading angle, deg East of North 77-55*H 77- 6399

-1.8

-00:01.8

0.0260

0.0597

-47

-0.01

-2.7

- 7 - 0

-0.0468

.0070

1.0

00:01

0.0556

-0.087!*

-99

-0.02

5-2

- 7 - 0

-o. 01*87

0.0858

173



TABLE XXI.- COMPARISON OF PLANNED AND ACTUAL TRAJECTORY PARAMETERS - Concluded

Condition and,quantity Planned Actual Difference

Orbit parameters

Perigee altitude, statute miles . .

Perigee altitude, nautical miles

Apogee altitude, statute miles

• . Apogee altitude, nautical miles

. Period, mln:sec ...

Inclination angle, deg .

Maximum conditions

Altitude, statute miles . .

Altitude, nautical miles ' '. .

Space-fixed velocity, ft/s'ec ..... ~

Earth-fixed velocity, ft/sec

Exit acceleration, g units . . ....'........

Exit dynamic pressure, lb/sq ft

Entry deceleration, g units . . •.

•Entry dynamic pressure, Ib/sq ft . . .

Landing point

North latitude ." . -

Vest longitude

100.1

87.0

166.2

ll*.l*

88:32

32:52

100.03

86.92

162.17

146.92

88:29

32:5k

-0.02

-0.08

-4. OJ

-3.W

-00:03

0.02

166.2

1 1 * 1 * . 1 *

25,737.0

24,1*20.0

7-7

a966

b378

7.6

162.17

140.92

25,732.0

21*, 415.0

7-7

982

7-7

1*72

- 1 * . 0 3

- 3 . 1 * 8

-5-0

-5.0

0

16.0

0.1

25

21°07'N

68°00'¥

21°26'N

C68°1*1'W

00-19'H

'00 "1*1'W

aBased on atmosphere at Cape Canaveral. •• ,.

Based on 1959 AKDC model atmosphere. . - . ' . • . '

"Actual" landing coordinates shown atove were those resulting from the trajectoryintegration. The retrieval point 20 minutes after landing was reported as 21"25-6'N. and68"36.5'W- by the recovery ship.





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MERCURY NETWORK PERFORMANCE

The Mercury Network consists of the Mercury Control Center (MCC)atCape Canaveral (CNV); stations at the Atlantic Missile Range (AMR), Bermuda

(EDA), and at fourteen other locations along the orbital ground trackj andcommunications and computing centers at the NASA Goddard Space Flight Center.The network affords a data acquisition capability for real-time monitoring,mission control, and postflight analysis. This section presents informationon the performance of the Mercury Network in the areas of communications,telemetry, tracking, computing, and command systems. .A brief descriptionof the Mercury Network is given in reference' 12, and appendix A of referenceamplifies .the performance analysis presented herein.

Mercury Network performance was excellent, all systems were fullyoperational at lift-off, and the few minor malfunctions which occurred

did not affect the flight monitoring and control of the mission. Acqui-sition of data from tracking, telemetry, and air-to-ground (A/G)voicesystems was satisfactory in both quantity and quality for real-timemonitoring and postflight analysis. The relaying of A/G voice back to theMercury Control Center from all network sites with a point-to-point voicecapability contributed substantially to the real-time monitoring .of themission.

Tracking

Radar tracking on this flight was satisfactory and superior to that ofthe two unmanned orbital missions (MA- and MA-5)• All stations provideddata for all passes whenever the spacecraft was above their horizon. Thequantity and quality of these data were more than adequate. Minor problemsexisted in S-band phasing and handover, but these problems resulted in onlya negligible loss of data. The communications used for the phasing and .handover were very satisfactory. . ,

Interference of an unknown source caused some .concern on C-band atCape Canaveral and Bermuda, but this did not cause an extensive loss ofdata. It is apparent that the extensive maintenance, training, and refine-ment of tracking procedures for the network have yielded dividends. Satis-factory C-band tracking was accomplished during most of the "blackout"

period. Two Cape Canaveral radars had satisfactory S-band tracking forthe first 2 minutes of blackout, and they were then turned off because theend of their range interval had been reached. S- and C-band radar-trackingcoverage is shown in figures 62 and 63, respectively. The performance of theacquisition aid unit was satisfactory, and radar acquisition, in all cases,was accomplished without difficulty.

Data Transmission

The transmission of both high-speed and low-speed data was satisfactorythroughout the mission.



Trajectory Computation

At lift-off, the selected source for display at the MCC was the outputof the IP 7090. The FPS-16 tracking at Cape Canaveral was utilized untilapproximately 00:00:45, at which time the IP 7090 switched to Azusa trackingand these data were displayed for approximately the next 20 seconds. Launch-

vehicle guidance data through Goddard Space Flight Center were then selectedand displayed throughout the remainder of powered flight. The guidance-system radar acquired both rate and track at 00:01:08, and the systemlocked-on throughout the remainder of the powered phase. The quality of theguidance system radar data was excellent up to sustainer engine cutoff (SECO)and during the go-no-go computation.

The programed phase of the flight showed minor deviations of 0.75° inflight-path angle and 1.0 nautical mile in altitude at booster engine cut-off (BECO). After launch-vehicle staging, the radio guidance system correctedthese deviations. A maximum deviation of 2.8 nautical miles in crossrange

and an apparent residual of more than 50 ft/sec in yaw velocity was indicatedat sustainer engine cutoff. The yaw velocity was nominal up to approximately55 seconds before cutoff, but it then appeared to lack response to steeringwith the final results as stated previously. The calculated landing pointat the Canary Islands station, however, was right on the expected groundtrack, which was difficult to resolve with a residual velocity of 50 ft/secin yaw. It was later disclosed$ however, that an error in the scale of theplotboard at MCC existed, which explains the velocity deviation and accountsfor this misrepresentation. The cutoff conditions displayed in MCC arelisted in table XXII.

Low-speed tracking data from the remote sites were excellent; hencethe orbit was well defined by the end of the first orbital pass. Subsequenttracking during the second and third passes showed negligible improvementin the orbit parameters. The number of radar observations received fromeach site is shown in table XXIII.

The primary computer was lost during the second pass between Hawaiiand Point Arguello, California. A restart was made in less than 5 minutesusing the Hawaii vector; thus, the computer was ready to accept the White Sands,N. Mex., data. Because of a malfunction of the secondary computer, datafrom Corpus Christi, Tex., and Eglin Air Force Base, Fla., for the secondpass and data from Eglin Air Force Base on reentry were ignored.

During reentry, tracking data appeared to pinpoint the landing locationwith a high degree of confidence. The final values from the Goddard computersindicated a difference of only 2 nautical miles between the landing locationas obtained from the California Station data and the Cape Canaveral FPS-16data (seetable XXIV). However, the landing point reported by the recoveryship, as well as that computed by the Cape IP 7090 computer using CapeCanaveral and San Salvador FPS-16 data, did not agree with the Goddard computa-tions . The actual landing point was 39 nautical miles short of the planned

196





was not as satisfactory as the UHF. The HF voice system was of particularvalue during the first and third passes when the stations at the Canary Islands,Point Arguello, Calif., Guaymas, Mexico, Zanzibar, Indian Ocean Ship, Muchea,Astralia, and Canton Island were able to converse with the astronaut beyondthe capability of the UHF system. It is interesting to note that, in someinstances where the HF was being used as the spacecraft approached the station,

the quality of communications improved considerably as the elevation anglebecame positive, particularly as the switch was made to UHF.

Figure 66 shows approximate coverage compared with times above thevisual horizon. .Through the air-ground voice system, MCC was able tofollow the recovery procedure and monitor all conversations until after thespacecraft was aboard the recovery ship. •

Command System

The command system for MA-6operated in a satisfactory manner during

the mission. The few airborne-system anomalies are discussed below. The600-watt stations appeared to have had coverage beginning at a slant rangeof ^00 to ^50 nautical miles, and the 10-kilowatt stations had coveragebeginning at a slant range of 650 to 700 nautical miles. . A summary of thecommand handover exercises is shown in table XXVI and a summary of thecommand transmissions is shown in table XXVII.

Ground system.- There were several, problems involving the commandequipment and the coder relay panels during the month prior to launch;however, no delays in the launch countdown resulted. A total of 11 functionswas successfully transmitted from the sites: auxiliary sustainer cutoff(ASCO) was transmitted from San Salvador, three sets each of R and Z cali-

brations were transmitted from Muchea, and two sets each of R and Z calibrationswere transmitted from Cape Canaveral. Command coverage from all sites wassatisfactory with the exception of Muchea on the third pass. A combinationof slant ranges in excess of ^50 nautical miles, airborne antenna patterns,and 600 watts of RF power resulted in only 1 minute and 30 seconds ofcoverage above receiver threshold.

Airborne system.- Command receiver "A," operating from the l8-voltisolated bus,appeared to be much more sensitive to signal strengths above30 microvolts than receiver "B," which operates from the l8-volt standbybus. Below 30 microvolts the operation of both receivers coincided. The

onboard recorded signal strengths, although acceptable, were about 6 dbbelow those of the unmanned orbital missions (MA-4and MA-5). The airborneantenna pattern problem, which was experienced on MA-1+- and MA-5, was againevident from the M A ^ - 6 onboard records. Spacecraft attitude changes aredefinitely reflected on the signal-strength records. lonization blackouton the command frequency occurred between 0 : 3:03.5 and 04: 7:12, whichis a period of 4 minutes and 8 seconds.

Triggering of the "All-Function Events Channel" occurred five timesduring ionization blackout. The tone channels triggered are unknown, but

198



are coincident with a burst of signal into the receivers. It is known thatthe tones keyed were not clock changes, R and Z calibrations," nor a Maydaysignal. Postflight tests of the communication system revealed that inter-ference between the telemetry and UHF voice transmitters produced a signalwith a frequency on the edge of the command-receiver bandwidth. An increaseof 0.5 me in the assigned low-link telemetry frequency could correct this.

The characteristics of the inputs to the command receivers are shown as anoscillograph-record reproduction in figure 67.

What is assumed to be random noise with a signal strength from 1 to4 microvolts was recorded between 01:1 :00 and 01:15:07. Command carrier wasnot present during this period.

Ground Communications

All the ground communication networks provided good support for themission. Except for a few short prelaunch outages, all voice, teletype,

and datalines were available at all times, and the quality of transmissionwas satisfactory. Single-sideband voice communication with the two shipswas very satisfactory, as provided by AMR. Part of the link from the IndianOcean Ship had to be relayed through Ascension Island.

199



TABLE XXII.- ORBITAL INSERTION CONDITIONS DISPLAYED AT MCC

Quantity Actual Nominal

Velocity ratio with posigrades . . . .

Space-fixed flight-path angle,

deg

Insertion altitude, nauticalmiles

Inclination angle, deg

Orbit duration capability , number ofpasses

Insertion velocity withposigrades, ft/sec

Apogee altitude, nautical miles . . .

1.0002

a

-0.0674

32.540

7

25,728

141

1.00058

0

87

32-52

25,737

T?hese values represent the average of the go-no-go computations.

Only the number ascertained by computation; actual orbitcapability was approximately 100 passes.

200



TABLE XXXIX. - SUMMARY OF LOW-SPEED TRACKING DATA

[_Verlort not used in differential correction after correction onFP6-16J

Facility

Station Radar

Total possible

valid observations

Total observations

made

Valid

observations

Nonvalid

observations

Differential

correction

First pass

Bermuda

Bermuda

Canary Islands

Muchea ,

Australia

Wooniera,

Australia

Havaii

Havali

Point Arguello,

Calif.

Point Arguello,

Calif.

Guaymas,

Mexico

White Sands,

H . Hex.

Corpus Christ!,Tex.

Eglln Air Force

Baae, Fla.

Eglin Air Force

Baae, Fla.

Cape Canaveral,

Fla.

FPS-16

Terlort

Verlort

Verlort

FPS-16

FPS-16

Verlort

FPS-16

Verlort

Verlort

FPS-16

Verlort

FPS-16

Verlort

FPS-16

71

7168

82

40(a)

(a)

(a)

(a)

65

34

6k

40

(b)

566172

93

85

6k

56

71

39

71

53

5365

76

40

51

28

46

38

43

38

9

17

45

13

28

25

1

28

43

50

50

40

I*

28

0

38-

35

Second pass

Bermuda

Bermuda

Canary Islanda

Muchea,

Australia

Voomera ,

Australia

Hawaii

Hawaii

Point Arguello,

Calif.

Point Arguello,

Calif.

Guaymas,

Mexico

White Sands,

H . Hex.

Corpus Christi ,

Tex.

Eglin Air Force

Base, Fla.

Eglin Air Force

Base, Fla.

Cape Canaveral,

Fla.

FPS-16

Verlort

Verlort

Verlort

FPS-16

FPS-16

Verlort

FPS-16

Verlort

Verlort

FPS-16

Verlort

FPS-16

Verlort

FPS-16

66

(b)

54

80

33

15

56

38

(b )

( < = )

in

60

1.0

(b )

56

54

61

6515

34

29

63

64

42

58

47

46

60

291530

28

31

60

41

54

9

9

1

3604

1

32

!>

1

ll

42

34

50

29

15

28

31

• 4 7

38

Uli

Third pas

Bermuda

Bermuda

Canary Islands

Muchea,

Australia

Woomera ,

Australia

Hawaii

Havaii

Point Arguello,

Calif.

Point Arguello,

Calif.

Guaymas,

Mexico

White Sands,

H . Hex.

Corpus Christi ,

Tex.

Eglin Air Force

Base, Fla.

Eglin Air Force

Base, Fla.

Cape Canaveral,

Fla. .

FP3-16

Verlort

Verlort

Verlort

FPS-16

FPS-16

Verlort

FPS-16

Verlort

Verlort

FPS-16

Verlort

FPS-16

Verlort

FPS-16

65(b)

(a)

70

(a )

64

Il2

(b)

(c)

41

53

40

(b)

60

69

38

23

20

1

61

60

23

33

58

65

3814

17

34

54

22

27

2

< l

0

9

3

27

6

1

6

49

50

38

17

34

1*1

18

22

Out of range

Data not available

CPassive

201



l» • •• •

- TABLE XXIV.-7 SUMMARY OF LANDING-POINT PREDICTIONS

BASED ON RADAR TRACKING

ConditionNorth

latitudeWest

longitude

Based on Point Arguello, Calif. ,retrofire time

Differential correction based on

Point Arguello, Calif., tracking

White Sands, N. Mex., tracking

Corpus Christi, Tex., tracking

Eglin Air Force Base, Fla., tracking

Cape Canaveral, Fla., tracking

Cape IP 7090 landing data:

Based on Cape Canaveral,Fla.,FPS-16 tracking

Based on San Salvadore FPS-16

tracking

Value reported by the recovery ship

21°08'

21°32'

21°29'

21°29'

° '21°25

68°01'

68°55'

68°53'

68°53T

68°3l'

68°35'

68°37'

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• •••• •• ••

TABLE XXV.- TELEMETRY RECEPTION SUMMAH - Concluded

(c) Third orbital pass

Station

Cape Canaveral,Fla.

Bermuda

Atlantic OceanShip

Canary Islands

Kano, Nigeria

Zanzibar

Indian OceanShip

Muchea ,Australia

Woomera ,Australia

CantonIslands

H a wa ii

Point Arguello,

Calif.

Guaymas,Mexico

Corpus Christi ,

Tex.

Eglln Air ForceBase, Fla.

Cape Canaveral,Fla.

Cape Canaveral,Fla.

Telemetry

Acquisition,hr:min:sec

03:06:51

03:09:56

03:24:44

(b)

(b)

(b)

03:>*6:55

03:56:31

04:03:16

(c)

04:21:49

04:31:17

04:33:44

04:36:53

04:39:00

04:40:52

Loss of signal,hr:min:sec

03:13:46

03:17:03

03:32:25

03:56:49

04:04:12

04:06:19

04:28:49

04:37:57

04:39:49

04:42:32

04:42:52

04:42:55

Dec om m utator

Acquisition,hr:min:sec

03:06:53

03:10:06

03:25:06

03:48:10

03:56:49

04:03:31

04:22:02

04:31:27

04:34:04

04:36:58

04:39:21

04:40:56

d04:47:22

Loss of signal,hr:min:sec

a03:15:42

03:17:03

03:31:22

03:56:30

04:04:08

04:06:01

04:23:39

04:37:56

04:39:39

04:42:34

04:42:48

04:42:53

ÔlnSOilO

Slant range

Acquisition,nautical mile

780

870

900

1,050

1,030

870

920

900

770

930

800

590

Loss of signal,nautical milef

920

900

920

1,100

94 0

1,000

770

540

740

603

500

150

Elevation

Acquisition,deg

0

-1.2

-.5

-1

-.7

1

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-2

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-3

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j.

Loss of signal,deg

0

-1.2

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-1.4

-.16

-1.4

.1

3-6

-1.1

-5

1.4

17

Includes data from down-range stations via submarine cable

Out of range

°Not applicable

from down-range stations via submarine cable

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221



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Figure 6j.- Concluded.

High frequency

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Low frequency

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223



CONCLUDING REMARKS

The significance of the first United States manned orbital flightis well recognized throughout the world. This and other papers haveemphasized the unqualified success of the MA.-6 mission. All spacecraft,

launch-vehicle, and Mercury network systems performed satisfactorily towardthe fulfillment of the specified primary test objectives.

The problems,, malfunctions, and anomalies which did occur were notsufficient to compromise the mission. The three occurrences which causedthe greatest concern during the flight were: the thruster failure in thereaction control system, the erroneous signal that the heat shield hadreleased prematurely, and the spacecraft oscillations and early drogueparachute deployment experienced during reentry. Each problem has beenintensely investigated and appropriate action taken for future missions....

Regarding the reaction control system failure, evidence exists whichindicates that small metallic particles from the fuel distribution screenshad lodged in the orifices of the 1-pound yaw thrusters and may have causedthe sporadic thruster operation. Although other explanations seem plausible,the cause of the control system failure is not definitely known. Testingin this regard is continuing.

The improper signal during the orbital phase of the mission that theheat shield had released has been conclusively traced to a faulty limitswitch. A-decision to retain the retropackage during reentry, which wouldprevent the heat shield from parting the spacecraft prematurely, was madeby the Operations Director at the Mercury Control Center and this decisionproved to be sound even though it was-of no consequence. Immediate stepshave been taken to improve the production techniques and quality control ofthese switches, as well as to incorporate necessary design modifications.In addition, the appropriate instrumentation circuitry has been revised tomake more effective use of redundant elements. The series of oscillationsexperienced early in the phase after reentry were adequately reduced in theauxiliary damping mode, which employs one automatic portion of the reactioncontrol system. However, a greater-than-expected rate of fuel usage duringthe orbital phase resulted in depletion of the automatic fuel during thereentry period and oscillations again diverged. The astronaut attempted tocontrol the oscillations manually, but human reaction time at the frequencies

experienced and the remaining manual fuel were insufficient to control thesituation adequately. The oscillations continued to diverge until thepremature drogue-parachute deployment, at which time the oscillations werereduced to nearly zero. The cause of the early drogue signal is unknownbut a stray electrical signal in the sequential circuit may have been re-ceived by the drogue mortar. Postflight tests did not reveal any parachutesystem anomalies.

Probably the single most important result of the MA-6flight was theproof that man not only can successfully function in terrestrial space,but that he can perform effectively under urgent conditions. Comprehensive

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medical examinations before and after the flight disclosed no adverse physio-

logical and psychological effects as a result of ^r- hours in a space environ-

ment under weightless conditions.

The success of the MA.-6 mission established another significant milestone

in the Mercury flight program, as did the two manned suborbital flights(MR-J and M R - i f - ) . The knowledge gained as a result of the noncritical problemswhich arose during the flight will be applied to spacecraft of subsequentMercury missions in a continuing program of system improvement. Possibly ofgreater importance, however, is the valuable experience which was achieved inthe area of space-flight operations, including the detailed checkout and pre-launch testing of the spacecraft, the real-time flight monitoring by personnelstationed around the world, and the vast recovery task which employed manymilitary vehicles and a staff of thousands. Therefore, this total missionexperience, of which the inflight results are but a portion, will enhance andlend confidence to future Mercury missions with more demanding overall objec-

tives.

Manned Spacecraft CenterNational Aeronautics and Space Administration

Houston, Texas, February 20, 1963.

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Documents

First United States Manned Three-pass Orbital Mission Mercury-Atlas 6, Spacecraft 13 Part 1 - Description and Performance Analysis