By JOHN H. BOYNTON, Mercury Project Office, NASA Manned Spacecraft Center; E. M. FIELDS, Chief Project Engineering Office, Mercury Project Office, NASA Manned Spacecraft Center, and DONALD F. HUGHES, Crew System Division, NASA Manned Spacecraft Center




[39] Project Mercury beg an in 1958 with some basic systems research and a number of feasibility studies to determine if a spacecraft could be built which would sustain man in orbital space and return him safely to earth. Although it was recognized that some system development would be required, many of the spacecraft systems could be synthesized from existing hardware. A top priority was placed on the spacecraft production from the contract award in 1959, and 3 years later Astronaut John H. Glenn, Jr., completed three orbital passes about the earth. In this time span, design, development, and qualification of the spacecraft and its systems were accomplished nearly concurrently. The ground and flight- test programs, which included hundreds of wind-tunnel tests and parachute drops from aircraft, provided an opportunity to develop flight systems and acquire operational experience as the program progressed. Though a continuing attention to engineering detail by technical specialists and management personnel throughout the project, the spacecraft and its systems were qualified for suborbital flight in approximately 2 years from the spacecraft contract award date. Many lessons have been learned which were not only applied to Mercury systems development, but which have been applied in more advanced space projects. Interesting conclusions regarding system performance can be derived by reviewing all of the flight results. The spacecraft control system was a source of considerable trouble during the project. However, when inflight failures of this type occurred, it was the backup capability of the pilot which made possible the successful completion of the mission. In fact, the pilot's ability to control accurately the spacecraft attitude was instrumental in three of the four manned orbital flights in completing the mission successfully when a malfunction was present in the automatic system. One of these control-system malfunctions, an electrical anomaly during Astronaut Cooper's mission and the only one of major significance in the spacecraft throughout the entire 34-hour flight, was successfully circumvented by the pilot's manual control during the critical retrofire and reentry maneuvers.




The initial goal of Project Mercury was to place a man into orbit successfully and return him safely to earth, and this objective was fulfilled in February 1962 by the flight of Astronaut John H. Glenn, Jr. This objective was confirmed 3 months later by the flight of Astronaut M. Scott Carpenter. The final two missions in Mercury constituted a continuation of a program to acquire new knowledge and operational experience in manned orbital space flight. The ninth Mercury-Atlas mission (MA-9) was planned for up to 22 orbital passes and was the concluding flight in the United States' first manned space program. The primary objectives of the MA-9 mission were to evaluate the effects on the astronaut of approximately 1 day in orbital flight, to verify that man can function as a primary operating system of the spacecraft; and to evaluate the combined performance of the astronaut and the spacecraft, which was specifically modified for the 1-day mission.


The MA-9 spacecraft, Faith 7, used by Astronaut Cooper in successfully performing the fourth United States manned orbital mission was basically similar to those used for previous orbital flights. The major exceptions were system modifications prompted by the extended nature of the mission, and these changes will be [40] discussed in later paragraphs. It is important to note, however, that since the original design of the Mercury spacecraft all major system concepts have remained essentially unaltered. Although some design and early development were conducted prior to the official award of the prime contract, the Mercury spacecraft was developed, qualified, and met its original objective of manned orbital flight 3 years after the spacecraft contract award in 1959. In this brief span of time, many lessons have been learned and much experience has been gained in the design, development, and operation of manned orbital flight systems. In this paper, the intent is to describe briefly the original design philosophy, discuss the system development and qualification experiences, and present a summary of the experiences relating to systems performance.


Design Philosophy


In the initial design of the Mercury spacecraft, two guidelines were firmly established: (1) to use existing technology and off-the-shelf equipment wherever practical and (2) to follow the simplest and most reliable approach to system design. These guidelines were administered to provide for the most expedient realization of program objectives. The original Mercury concept also included a number of mandatory design requirements which were imposed on the spacecraft contractor:

(1) The spacecraft must be fitted with a reliable launch-escape system which would rapidly separate the spacecraft with its crew from the launch vehicle in case of an imminent disaster.

(2) The mode of reentry into the earth's atmosphere would be by drag braking only.

(3) The spacecraft must carry a retrorocket system capable of providing the necessary impulse to bring the vehicle out of orbit.

(4) The spacecraft design should place prime emphasis on the water- landing approach.

(5) The pilot must be given the capability of manually controlling spacecraft attitude.


In many design areas, there existed no previous experience in reliable system operation which could be applied to the Mercury concept, and new development programs had to be initiated. In addition, there was no information pertaining to man's capability to operate under space environmental conditions, particularly environmental weightlessness; therefore, all of the spacecraft systems which relate to crew recovery from orbit had to be designed for automatic operation and many had to include redundancy. It has since been learned that man is not only a contributory element but a necessary part of the spacecraft. It is important to note that because of the pilots demonstrated ability to function as a primary operating system of the spacecraft, some of the redundant elements were not required and were deleted.


The spacecraft systems (fig. 3-1) include the heat protection, mechanical and pyrotechnic spacecraft control, communications, instrumentation, life support, and electrical and sequential systems. The mechanical and pyrotechnic system group comprises the separation devices, the rocket motors, the landing system and the internal spacecraft structure. These systems have been described in previous literature (refs. 1 to 10); therefore, detailed descriptions are not included in this paper.


The design requirements stated earlier involved certain implications for these systems The launch-escape system was found to be most practical if it incorporated a solid rocket motor to propel the spacecraft rapidly away from the launch vehicle during an abort in the atmosphere. This type of system needed to provide a high level of thrust for a brief time period should be easily handled in the field and should require a minimum of servicing. The tower arrangement could be readily assembled to the spacecraft and jettisoned during powered flight once it no longer was required for abort.


An important design feature of the Mercury spacecraft was the favorable manner in which the astronaut was exposed to flight accelerations. For all major g-loads, which occur during powered flight, launch-escape motor thrusting, posigrade motor thrusting, retrograde motor thrusting, reentry, parachute deployment and touchdown, the pilot experienced acceleration in the most favorable manner, one that forces him into the couch (fig. 3-2).


The mode of reentry was specified to be drag braking only because of simplicity. This concept implied that the configuration should be a blunt body with high drag properties having a slender afterbody, primary because of heating considerations. Thus the bell shaped Mercury [41] configuration was evolved, and the heat-protection system was devised to accommodate this shape. Originally, a beryllium thermal shield employing the heat-sink principle was specified. The specification was later changed to provide a more efficient ablation-type heat shield, which was used on all Mercury- Atlas orbital missions. Because the heat flux was expected to be considerably less on the afterbody than at the heat shield, a combination of insulation and thin shingles constructed of an alloy to withstand high temperature was calculated to be sufficient in maintaining the temperature of the pressure vessel at a safe level. The exterior finish of the spacecraft body was intentionally made a dull black because of its high emissivity and, therefore, favorable thermal radiation properties.


Again, because of their reliability and ease of handling and servicing, solid propellants were chosen for the retrorocket system. For even greater reliability, however, a system of three solid rocket motors, any two of which would effect a safe reentry, was chosen. These three rocket motors, together with three additional rockets to effect spacecraft- launch-vehicle separation, were assembled in a jettisonable package to permit a clean reentry configuration.


For the period during and after touchdown, the spacecraft had to meet two basic requirements. These requirements were: ( a ) the structure had not only to retain its integrity such that it would be habitable after landing and (b) the touchdown decelerations had to be reduced to an acceptable level. Touchdown deceleration was primarily limited by the human tolerance to acceleration; and, because of the blunt shape of the spacecraft, touchdown decelerations of as high as 50g could have resulted even for a water landing. Therefore, a landing-shock attenuation system was designed which consisted of a fiberglass fabric bag with holes in it and was attached between the spacecraft structure and the ablation shield. Prior to landing, the ablation shield would be deployed and the shield weight would extend the bag, which would fill with air and provide a cushion against the landing shock. The landing bag arrangement adequately attenuated the landing deceleration loads to approximately 15g.


In addition to the automatic and rate control modes of the attitude control system, two manual control modes, one electrical and the other mechanical, were provided the astronaut. This control-mode arrangement had the feature that,


Figure 3-1. Spacecraft interior arrangement.


[42] in the event of a spacecraft power failure, the direct-linkage mechanical mode would still be available for control. The two manual control modes were each supplied control-system fuel from separate tanks for additional reliability. Although the thrust units were designed to provide an impulse sufficient for the majority of spacecraft maneuvers, these redundant manual control modes could be used simultaneously, if desired, in critical situations, such as retrofire and reentry, where rapid response to undesirable attitude rates might become necessary.


A monopropellant reaction control system using hydrogen peroxide as the fuel was chosen for the spacecraft control system to provide the simplest system design and installation. Furthermore, similar systems had already been developed for use on other space vehicles. A flexible bladder under pressure provided a positive means of fuel expulsion.


Many challenging design problems were encountered in the remaining spacecraft system because of the new operating environment. As a result of the need to provide flight-control support on the ground, the requirement for multiple redundancy and high reliability in the communications system was evident. Although part of the instrumentation system was not required for flight safety and mission success, certain parameters, such as those which indicate the physiological well-being of the crew and the propel operation of critical spacecraft systems, were necessary for effective flight control and monitoring. The remainder of the instrumentation data was acquired to complement the flight-control parameters for use in postflight


Figure 3-2. Acceleration loading for various flight phases.


[43] analyses of system performance. New design areas were opened up in the fields of gas partial pressure measurement and of bioinstrumentation, such as long term attachment of human sensor leads. The life-support- system design considerations involved a development task since it was concerned with the sustenance of the astronaut and his protection from the hard vacuum of space, as well as from the widely varying temperature conditions associated with an orbital-flight profile. This system also was required to provide for the management of the cooling and drinking water in the spacecraft, the food to be consumed by the pilot, and his normal liquid wastes, again in the weightless environment. Although pressure suits and cooling equipment had been used in high-performance aircraft, only art of this experience could be directly applied to the design of the Mercury environmental control system because of weightless flight and more demanding performance requirements. In the electrical and sequential design area, the application of previous design work and use of off-the-shelf components was made. But the very nature of the mission and the requirement for reliability' automation, and system redundancy imposed a degree of complexity somewhat greater than any previous manned flight system. This increased electronic complexity in turn, made it more difficult to insure interface compatibility, eliminate stray voltages (back-door circuits), and minimize system sensitivity to current transients.


As an example of the consequences of stray voltages, the Little Joe-1 mission, the first launch attempt using a full-scale Mercury spacecraft, is cited. This test, conducted at Wallops Island, Va., was in the final moments of countdown when, during a spacecraft battery charging operation, a stray voltage initiated the launch escape sequence. The spacecraft was separated by the escape motor from the launch vehicle, and the drogue parachute was properly deployed. Because the battery had been only partially charged, sufficient current was not available to deploy the main parachute, and the spacecraft was destroyed upon landing. This back-door circuit was subsequently located and eliminated.


Because of work conducted immediately prior to and in the early period following contract award, the system-design phase of the project proceeded at a rapid pace. Wind-tunnel research, studies by prospective subcontractors and vendors, the joint participation of key NASA and other government installations, and early design studies by the eventual prime contractor all helped to facilitate the design effort and make possible the early availability of test hardware.


Based on the total Mercury experience, one of the underlying principles during the initial design period should be an emphasis placed on ''designing for operation." For example, one of the lessons learned was that the instrumentation system should be designed with mission flexibility as a guide, such that, in the later phases of the program, new instrumentation requirements can be handled with a minimum of complication. In still another area, it was learned that component accessibility can be extremely important where schedule demands become critical. Certain time-critical systems and short-life components must be easily accessible in order to minimize the degree of disturbance to other systems and the time required to replace these types of units. Because of the weight and volume constraints, this concept could not faithfully be applied in the design evolution of the Mercury spacecraft, and significant penalties have been experienced wherever items needed to be removed under a tight schedule. It was learned in Mercury that all systems requiring manual operation by the astronaut must be designed with the limitations of the cabin volume (see fig. 3-3), suit mobility, and weightlessness in mind.


Development and Qualification


As in any development program, one of the original ground rules at the outset of Project Mercury was to conduct a logical and progressive test program. This concept was closely maintained from the beginning, of the project through the flight of Astronaut Cooper last May. Success in certain phases of this test progression has made possible the elimination of certain backup or follow-on flights. Since the time that Mercury was initially conceived, literally thousands of individual tests have been conducted in which test articles were used in all forms from components to full-scale spacecraft and under all combinations of real and [44] simulated operating conditions. For example, during the 1-year period from November 1959. about 10 months after the prime contract was awarded, to November 1960, some 270 hours were spent in testing the environmental control system in the altitude chamber, with a man wearing a pressure suit in the chamber to load the circuit more realistically. Early in 1961, further tests were conducted, often using astronauts, in the centrifuge to qualify the environmental system under acceleration loads.


For convenience, the spacecraft-system testing can be grouped into ground tests and flight tests of special test articles. The ground tests, in turn, can be categorized into areas of research, design, development, qualification, acceptance, and checkout. The discussion of development flight tests, which will be restricted to those using other than production spacecraft, consists of research studies, development tests, and qualification programs. The performance of the production spacecraft will be discussed


astronaut seated in spacecraft

Figure 3-3. Photograph of spacecraft interior.


in a later section of this paper. It is interesting to note that because of the rapid pace dictated by the high priority of the program, many of the individual test programs were conducted concurrently. This technique involved some risk, since, had a major problem developed, the expense in both time and money could have been considerable. The following paragraphs relate the more salient lessons learned during the formal Mercury development and qualification test program.


Ground Testing


The research tests included those which attempted to verify design theories or sought new methods for accomplishing a given design task, whether it was a structural assembly, a heat-protection system, or improved methods of instrumenting the spacecraft and its crew. Hundreds of tests of this type, particularly those conducted in the wind tunnel, were [45] carried out in the early phases of the Mercury effort at many of the NASA centers and at the contractor's plant. These tests will always be required when a new flight spectrum in a relatively unknown operational environment is penetrated, as it was in Mercury. It was tests of this kind which established the basic Mercury configuration, a shape which has already been projected into more advanced manned space programs.


The design testing, exemplified by the breadboard layouts in the case of electrical and sequential circuitry, was conducted jointly by the NASA and the contractor. This effort made possible the proof testing concurrent with initial design studies. Many thousands of tests were conducted, such as those in the design of the spacecraft-control-system thrust chambers, once the initial concept had been established.


When the basic design concept had evolved to a working hardware item, development testing served to expose this concept in the laboratory to the many combinations of operational and environmental conditions expected in space. Development testing was naturally hampered by the fact that weightlessness, a prime example, could not be adequately simulated on the ground; and this very deficiency resulted in an ineffective design for the water separation device of the environmental control system. The development of Mercury systems was a continuing program through the final mission and was aimed at mission flexibility, even after the spacecraft had been basically qualified for manned orbital operation. It was during the development testing that facets of the design which pertain to all aspects of its use were most evident, including the design-for-operation standards. It is in this testing area that engineering mock-ups have proved to be extremely valuable. In the case of the landing system, drop tests of boilerplate spacecraft were made to develop the landing-system deployment sequence and operation. Tests were made in the altitude chamber to verify that systems could operate for their required life cycle under realistic conditions. In essence, the development- test phase provided a means of validating the design concept and proving its intended reliability features.


Qualification testing conducted on the ground can further impose realistic operational conditions on a test article in various combinations for the specific purpose of verifying its reliable operation for inclusion as a final flight article. That is, there can and should be more shall one type of qualification program for a given component, subsystem, or system, but these programs should become progressively more demanding on the capability of the hardware. In this testing, area, adherence to prescribed test criteria must be rigorously enforced. The various combinations of qualification tests can be grouped into environmental tests, load tests, and performance tests with each of these groups having a specific purpose. Sometimes, the test conditions are not realistic enough or are not sufficiently demanding to reveal system weaknesses. During Mercury, for some of the subsystems, it was not until the actual unmanned flights that a system could be fully qualified for manned operation. For example, the launch-escape tower was subjected to all expected environmental conditions, all exhaustive series of load tests, and the operational situations associated with the launch- escape-system performance tests. Yet in the actual qualification flight program the heating loads on the truss structure of the tower were found to be more critical than had been calculated. Ground qualification is relatively inexpensive compared with full-scale flight qualification, and any system discrepancies which can be revealed in this phase will yield rewards in terms of time and expenditures later on. For example, during an early qualification test, it was found that the original igniters in the retrorocket motors would sometimes fail and blow out through the rocket nozzle before the main propellant grain had been ignited. New igniters, actually miniature solid rockets, were substituted for the original igniters. Had this system characteristic been overlooked through the manned orbital flights, the consequences could have been catastrophic. For flight-acceptance tests on units scheduled to be installed in flight vehicles, however, it was found that care should be taken not to over-test the article to the point at which its useful lifetime is approached or exceeded. During qualification testing, one must be assured that the unit being tested is not a "handmade" article and that, later on, a similar production version will not fail because it does not have the same ability to withstand the [46] testing environment. Of course, a critical requirement for the qualification program is that the test conditions imposed on the hardware exceed those expected to be present in the design environment in order to provide a safe margin for manufacturing deviations and unanticipated design weaknesses. It was found in Mercury that no single qualification criterion necessarily applies to all systems, and local operating conditions must be evaluated specific ally for each system to insure that they are adequately accounted for in the qualification test environment.


It was learned in Mercury that, whenever a significant design change is to be incorporated into the spacecraft, a new hardware qualification program should be initiated to requalify major systems. Approximately 1,000 hours of test time were accumulated on a full-scale spacecraft in a program called "Project Orbit" which was conducted in a vacuum-thermal facility to insure that, during the orbital flight program, systems would maintain their previously demonstrated performance. is an example, when the spacecraft thruster assemblies were modified as discussed in this paper, the modified assemblies were tested in a vacuum chamber as part of the Project Orbit testing. These tests, using hydrogen peroxide, were made to determine if exposure to combined temperatures and low pressures for the expected duration of the mission would have adverse effects on the operation of the thruster assemblies. It was found to be most effective if actual operating conditions and procedures, including preflight checkout tests, could be realistically simulated in order to expose hardware to a complete operating cycle. Since system qualification and operating reliability are so closely related, the reader is referred to the paper entitled "Reliability and Flight Safety" for additional details.


Finally, the acceptance and checkout tests which are conducted by using actual flight hardware involve the same recommendations previously mentioned, those of avoiding over testing, realistic operational test conditions, and thoroughness. It was learned in Mercury that, if tests of this type are conducted at multiple stations across the country by separate groups, the test procedures must be consistent if the test results are to be comparable. It is essential to repeat a system checkout if the system has been disturbed for any reason, such as the removal of another system where a definite interface exists. The acceptance and checkout aspect of ground testing is more thoroughly discussed in the paper entitled "Spacecraft Preflight Preparation."


Flight Testing


This brief discussion of the development flight phase of Mercury will be limited to those flights where specially configured test vehicle (boilerplate spacecraft) were employed. Because the experiences gained by flights of production spacecraft are of more operational significance, they will be presented in the next section, Systems Performance. The flight-test program began with a number of tests in which spacecraft models were flown by using small multistage rockets. These tests provided preliminary data on the aerodynamic properties of the chosen external configuration. Almost concurrently with these flights, tests of the parachute systems were staged in which boilerplate spacecraft were dropped from cargo aircraft. These "drop tests" were initiated as an important step in the early design and development of the landing system. Specifically, the drogue parachute was developed by utilizing a weighted pod, which was dropped from an aircraft at high altitude. Other early flight tests included off-the-pad, or beach, aborts to develop the launch- escape system. In 1959, a reentry flight was conducted in which a specially designed and instrumented spacecraft and an Atlas launch vehicle were used to provide aerodynamic-heating data in the real flight spectrum. This flight, termed "Big Joe, was the first test in Mercury in which the Atlas was used. It was as a result of the data derived during this flight that the shingles initially on the spacecraft cylindrical section were replaced with somewhat thicker shingles made of beryllium to provide for more effective heat protection. The final series of early flight tests used the solid-propellant Little Joe vehicle (shown in fig. 3-4) to test the launch-escape system concept at critical inflight abort condition. Although most of the early flight tests were of a developmental nature, their missions served to qualify critical flight systems for later, more demanding flight tests. The intermediate series of aircraft drop tests, for instance, was [47] completed to qualify the parachute and landing-shock attenuation systems. During this test phase in Mercury, valuable system improvements were incorporated at a minimum of cost and time.


rocket on launch pad

Figure 3-4. Mercury Little-Joe launch-vehicle configuration.


Weight Growth


A critical problem which was present throughout the Mercury program was that of weight growth. This problem, which seems to be characteristic of any development program where high performance and reliability are required, almost defies the steps taken to control weight. Figure 3-5 depicts the weight chronology of the spacecraft s orbital configuration. The maximum growth in weight was approximately 10 pounds per week in the very early phases of the program, but this figure was reduced to less than 2 pounds per week, or approximately 1/2 percent, at the final stage of the program. The launch weight of Astronaut Cooper's spacecraft, Faith 7, was some 700 pounds greater than the original design weight, despite repeated design reviews and other continuing weight controls. The lesson here is that proper planning must account for the inevitable weight growth in the design and development of high-performance spacecraft, since the consequences of not planning for it are either a degradation of the performance goals or exceeding the capability of the launch vehicle with its attendant delays.


Attention to Detail


One of the most significant lessons learned from the Mercury program was the need for a careful and continuing attention to quality and engineering detail in all phases of the program. The spacecraft is made up of many individual systems and components to form a complex entity, and only through a close monitoring of the design and development of each piece of hardware and its relationship to all other associated components is it possible to recognize and correct problems rapidly before a costly failure occurs. Many performance discrepancies could not be anticipated because of the lack of experience or the inability to simulate adequately realistic conditions in the early test program. Later tests, however, were established to reveal these anomalies with a minimum of cost and delay. Although somewhat limited by the lack of experience, attention to detail during the design phase resulted in the incorporation of


Figure 3-5. Weight chronology for Mercury specification spacecraft.


[48] system redundancy, where a direct relationship to mission success existed.


As a prime example of the attention given to the incorporation of redundancy in the detailed design of critical spacecraft components, the actuation system of the launch-escape-tower clamp ring w as backed up in nearly every component because of the serious consequences that would have resulted from a failure of the escape tower to jettison. In this system, the clamp ring is assembled at three points on its periphery, with each point being held by a dual explosive unit. Five of these six pyrotechnic units were ignited by an electrical squib, whereas the sixth was actuated by a percussion cap. Each of the electrical units incorporated a dual bridgewire. The automatic sequence was designed to send electrical signals from one power source to six of the bridgewires, with another but independent electrical supply for the remaining four bridgewires. Should the automatic relay fail, the astronaut was provided with a manual pull-ring which would energize the same jettison relay and also operate a gas generator to initiate the percussion cap, such that, in the event of failure in both the circuit to the sequencing relay and the two separate electrical power buses, the percussion cap would ignite. Actuation of any one of the six pyrotechnic explosive bolts was sufficient to effect proper separation of the escape tower from the spacecraft. The pyrotechnic circuit for the spacecraft-launch-vehicle adapter clamp ring was operated in a nearly identical manner.


During the development phase, an adherence to test specifications was maintained through a continued scrutiny of detailed performance results as they became available. Throughout the manned flights, attention to detail was necessary for an early recognition of possible problem areas, provided a means of responding to suggested action items, and precluded the occurrence of some system failures which ordinarily would have caused launch postponements and possibly a catastrophe.


Systems Performance


During the design of the Mercury spacecraft, one of the most important considerations was that, should individual components or even entire systems fail, some means would exist either to complete the mission safely or to conduct a successful mission abort so that crew safety would be maintained. A summary of the flight program objectives and results for the full scale spacecraft is given in table 3-1. Of primary significance in the table is the fact that during the manned flight phase, all major systems operated satisfactorily, although on three of these missions, the astronaut was required because of improper operation of the automatic control system, to conduct the retrofire maneuver manually. There were system malfunctions and performance discrepancies in each of these flights, but they were of such a nature that either a backup system or astronaut could circumvent the anomaly or that the failure of a component, such as an instrumentation sensor was not critical to mission success. The system experience during the flight program was characterized by a number of isolated component anomalies, rather than a critical failure of such magnitude that a catastrophe resulted. This system development, accounting for system malfunctions and performance discrepancies, the action taken to correct them, and the steps required to increase system capability for the extended flight of Astronaut Cooper, is discussed in the following paragraphs. Since system anomalies are discussed specifically as they pertain to the continuing development of the major spacecraft systems, references 5, 6, 8 and 10 should be consulted for a more detail performance discussion. Although random failures and system deficiencies are mentioned briefly herein, the greater emphasis is placed on system performance as it relates to design experience and the lessons which can be derived from actual operation of the systems in the space environment. Throughout the flight program, with the exception of the MA-9 mission, no changes were required specifically to accommodate a longer flight duration. The modifications made to the Faith 7 (MA-9) spacecraft including those incorporated to make possible the extended flight period are summarized in table 3-II. Each major spacecraft system will be discussed separately, as in previous reports on the individual manned flights (refs. 5, 8, an 10).


[49] Heat Protection System


The heat protection system performed satisfactorily throughout the entire program and essentially as designed.


Some cracking and slight delamination of the ablation heat shield following reentry have been experienced on certain flights, but this occurrence has been of no real consequence. It was established that this minor delamination did not occur during the reentry heating period and probably resulted from the shock sustained at landing Since the flotation attitude depends somewhat on the heat-shield weight, a slight modification was made to the Faith 7 spacecraft to provide for retention of any small portions which might possibly have broken away after touchdown. It has always been desirable to achieve the most upright position in the water to facilitate astronaut egress.


Temperature measurements were made at various depths in the ablation shields for the orbital flights, and the maximum values experienced during reentry are summarized in figure 3-6 for each flight. The measurements showed good agreement with predicted values and were satisfactory.


Mechanical and Pyrotechnic Systems


The mechanical and pyrotechnic systems consist of the separation devices, the landing system, the rocket motors, and the internal spacecraft structure. Each of the systems in this group is discussed separately.


There have been only minor problems with the separation devices. The primary separation planes (shown in fig. 3-7) are those between


Figure 3-6. Ablation shield maximum temperatures.


the launch-escape tower and the spacecraft cylindrical section, between the spacecraft and the launch vehicle, at the heat shield, and at the spacecraft hatch. In three of the earlier unmanned qualification flights, some difficulty was experienced in separating the spacecraft adapter umbilicals, but postflight examinations showed that the pyrotechnic charges ignited satisfactorily. Further investigation revealed, however, that aerodynamic loads during clamp-ring separation had caused the clamp-ring segments to damage the umbilicals. A minor redesign of the clamp-ring cover which protects these separation devices eliminated the problem. In the Mercury-Redstone 4 (MR-4) mission, the explosively actuated side hatch, incorporated for the first time for this flight, was prematurely released. The astronaut egressed rapidly through the open hatch, and the spacecraft subsequently took on sea water and sank before recovery could be effected. A postflight investigation involving a thorough analysis and exhaustive testing was conducted, but the cause of the malfunction has never been established. However, the landing and recovery procedures were altered for succeeding missions to minimize the possibility of this malfunction recurring. The only other performance anomaly with regard to separation devices occurred in the recent flight of Astronaut Cooper. Here,


Figure 3-7. Major spacecraft separation planes.

[50-51] Table 3-I. Mercury flight Program Summary.

Mission a

Spacecraft b

Launch date

Flight duration c, hr:min:sec


Basic test objective d

Summary of results e



Aug. 21, 1959



Max. dynamic pressure abort; evaluate launch escape and recovery systems.

Object. not met; inadvertent abort initiated during countdown.

Big Joe


Sept. 9, 1959



Ballistic flight; evaluate heat-protection concept, aerodynamic shape, and recovery system.




Oct. 4, 1959



Ballistic flight; qualify launch-vehicle structure; evaluate command system.




Nov. 4, 1959



Max. dynamic pressure abort; same as LJ-1.

Primary object. not met; escape motor ignition was late during reduced pressure region.



Dec. 4, 1959


Rhesus monkey

High-altitude abort; evaluate launch, abort, and recovery dynamics on S/C; recovery.




Jan. 21, 1960


Rhesus monkey

Max. dynamic pressure abort; same as LJ-1A; evaluate launch and abort.


Beach abort

S/C 1

May 9, 1960



Off-the-pad abort; qualify structure and launch escape system for simulated pad abort.

Successful; expended rocket motor and tower not separated as quickly as expected.


S/C 4

July 29, 1960



Ballistic flight; S/C-launch vehicle compatibility; thermal loads in critical abort.

Object. not met; mission failed at about 60 sec. after lift-off; S/C not recovered.


S/C 3

Nov. 8, 1960



Max. dynamic pressure abort; qualify launch escape system and structure.

Object. not met; S/C did not separate from launch vehicle.


S/C 2

Nov. 21, 1960


Simulated man

Suborbital flight; qualify S/C-launch-vehicle compatibility, posigrades, ASCS.

Test object. not met; launch vehicle shutdown at lift-off; S/C landing system correctly deployed.


S/C 2

Dec. 19, 1960


Simulated man

Suborbital flight; same as MR-1.

Successful; cutoff overspeed caused overshoot.


S/C 5

Feb. 21, 1961



Suborbital flight; qualify ECS, landing bag.

Successful; launch vehicle failed to shutdown until fuel depletion, S/C overshot by 130 miles.


S/C 6

Feb. 21, 1961



Ballistic flight; same as MA-1.



S/C 14

Mar. 18, 1961



Max. dynamic pressure abort; same as LJ-5.

Object. not met; escape rocket ignited early; S/C recovered intact.



Mar. 24, 1961



Suborbital flight; evaluate modifications to correct MR-1 and MR-2 malfunctions.



S/C 8

Apr. 25, 1961


Simulated man

One-pass orbital flight; evaluate all S/C systems, network, recovery forces.

Object. not met; launch vehicle failed to follow roll program; S/C escape system operated.


S/C 14A

Apr. 28, 1961



Max. dynamic pressure abort; same as LJ-5 and LJ-5A.



S/C 7

May 5, 1961


Alan B. Shepard

Suborbital flight; familiarize man with space flight; evaluate response and S/C control.

Successful; first American astronaut in space.


S/C 11

July 21, 1961


Virgil I. Grissom

Suborbital flight; same as MR-3.

Successful; premature hatch release caused S/C to take water and sink; astronaut recovered.


S/C 8A

Sept. 13, 1961


Simulated man.

One-pass orbital flight; same as MA-3.

Successful; open circuit in control system caused S/C to land 75 miles uprange; S/C recovered.


S/C 9

Nov. 29, 1961



Three-pass orbital flight; qualify all systems, network, for orbital flight recovery.

Successful; control system malfunction terminated flight after two passes.


S/C 13

Feb. 20, 1962


John H. Glenn Jr.

Three-pass orbital flight; evaluate effects and performance of astronaut in space; astronaut's evaluation of S/C and support.

Successful; first American to orbit earth; control system malfunction required manual retrofire and reentry; erroneous T/M signal, retropack retained through reentry; S/C landed 40 miles uprange.


S/C 18

May 24, 1962


M. Scott Carpenter

Three-pass orbital flight; same as MA-6; evaluate S/C modifications and network.

Successful; horizon scanner circuit malfunction required manual retrofire; yaw error caused S/C to land 250 miles downrange, recovery in 3 hours.


S/C 16

Oct. 3, 1962


Walter M. Schirra

Six-pass orbital flight; same as MA-6 and MA-7 except for extended duration.

Successful; partially blocked ECS coolant valve delayed stabilization suit temperature until 2nd pass; S/C landed 4 1/2 miles from primary recovery ship.


S/C 20

May 15, 1963


L. Gordon Cooper Jr.

Twenty-two pass orbital flight; evaluate effects on man up to 1 day in space; verify man as primary S/C system.

Successful; short circuit late in flight disabled ASCS, inverters, prompted manual retrofire and reentry; S/C landed 4 1/2 miles from ship.

a LJ- Little Joe launch vehicle mission; MA- Mercury-Atlas (launch vehicle) mission; MR- Mercury-Redstone (launch vehicle) mission; BD- Booster development.
b BP- Boilerplate spacecraft; S/C -spacecraft; S/C 10, 12, 15, 17 19 not used in flight program.
c Duration measured from lift-off to landing.
d ASCS- automatic stabilization and control system; ECS- Environmental control system.
e Object. - objectives of flight; prop. - propellant; T/M - telemetry.

[52] Table 3-II. Summary of Modifications to MA-9 spacecraft.




Spacecraft control system.

1. Removed rate control system (RSCS)

1. Not necessary; reduced weight by 12 lb

2. Added 15-pound-capacity fuel tank.

2. Additional control capability a

3. Installed modified 1- and 6-pound thrust chambers.

3. Improved reliability and operating characteristics.

4. Installed interconnect valve.

4. Improved control-fuel management.

Communications system.

1. Removed backup UHF voice transmitter.

1. Primary unit reliable, reduced weight by 3 lb.

2. Installed slow-scan television unit.

2. Inflight evaluation of TV for ground monitoring of astronaut and instruments.

Instrumentation system.

1. Deleted backup telemetry transmitter.

1. Primary unit reliable, reduced weight by 2 lb.

2. Changes recorder speed from 1 7/8 ips to 15/16 ips and programed.

2. Greater flight coverage necessary without changing recorder or reel size.

3. Deleted periscope.

3. Reduce weight by 76 lb; unnecessary for attitude reference.

4. Deleted low-level commutator

4. Served its purpose on previous flights.

Life support system.

1. Added 4 lb of breathing oxygen.

1. Necessary for extended mission

2. Installed parallel suit-coolant control valve.

2. Added reliability in case of partial valve blockage as experienced in MA-8

3. Added inline condensate trap

3. Existing condensate system believed ineffective

4. Added urine and condensate transfer systems with manual operation

4. Increase urine and condensate storage capability because of extended mission

5. Added 9 lb of cooling water

5. Increase cooling capability because of mission

6. Added 4.5 lb of drinking water

6. Necessary for increased mission duration

7. Added 0.8 lb of CO2 adsorber

7. Necessary for increased mission duration

Electrical and sequential systems.

1. Replace two 1,500 watt-hour batteries with two 3,000 watt-hour units

2. Necessary for extended flight duration.

2. Replaced two of three inverters

2. Improved thermal and operating properties.

a Tank intentionally serviced to only 10 lb. of fuel.


four of the five umbilicals, two between the spacecraft and the adapter and three between the spacecraft and the retropackage (fig. 3-8) failed to separate in a normal manner. Later analysis revealed that each of the malfunctioned disconnects (see fig. 3-9), which normally contained a dual charge came from a special test lot which did not contain the main charge of explosive powder. Somehow, this lot had been improperly marked as intended for flight hardware. The umbilical which separated normally contained the intended amount of explosive and came from a properly identified lot. The four umbilicals which failed to separate pyrotechnically were released through actuation of a backup mechanical device. This experience points up the necessity for close control of flight articles and a means for establishing that the hardware intended for flight satisfies prescribed specifications.


The landing system, which includes the main, reserve, and drogue- stabilization parachutes and the landing-shock attenuation system (landing bag), has never failed in flight during the production-spacecraft flight program. In the second Mercury-Redstone mission, the heat shield was lost after landing because the metal retaining straps and landing bag material to [53] which the shield was attached failed as a result of wave action and strengthening of existing straps for later spacecraft eliminated this problem. The only other anomalies in the operation of the landing system were concerned with the altitude of parachute deployment, and these. anomalies are discussed in the Electrical and Sequential Systems section. The successful performance of the landing system, particularly the parachutes, can be attributed to a thorough test program involving some 80 air drops of full-scale spacecraft.


spacecraft at launch site

Figure 3-8. Spacecraft photograph displaying retrorocket umbilicals.


The rocket motors include the launch-escape motor, the retrorockets, the posigrade rockets, and the launch-escape-tower jettison motor. All of the rocket motors used solid propellant, and their nominal thrust values are indicated in table 3-III. Each of these rocket systems has operated satisfactorily throughout the Mercury flight program. It was found early in the program the launch-escape tower did not separate rapidly enough from the spacecraft after an off-the-pad test because of thrust impingement on the tower; therefore, the tower-jettison rocket-nozzle configuration was subsequently changed from a one-to a three-nozzle arrangement. Because of reliable launch-vehicle operation, the launch-escape system was never needed for an atmospheric abort during the manned flight program, and the large escape motor successfully ignited each time when the system was normally jettisoned. An abort, however, occurred during the unmanned MA-3 mission, and the system operated satisfactorily.


diagram of umbilical

Figure 3-9. Schematic diagram of explosive umbilical disconnects.

Table 3-III. Nominal Rocket Motor Characteristics.

Rocket Motor

Number of motors

Nominal thrust each, lb

Approximate burning time each, sec





Tower jettison












The internal spacecraft structure has been compromised only once during a mission critical situation, a record which is essentially proved by the fact that water, following an ocean landing, had never entered the spacecraft in appreciable amounts, except in one instance, because of a structural failure. In the MR-2 mission following landing recontact of the heat shield with the large pressure bulkhead caused puncturing that resulted in a sizable leakage rate.


[54] The spacecraft was recovered, however, within a safe period. During postflight inspections of all manned spacecraft, some evidence of recontact by the heat shield upon landing has been present, but this damage to the large pressure bulkhead has been slight. The integrity of the spacecraft's load-carrying structure was especially proven during the Little Joe flight program. In one of these flights, the late ignition of one of the Little Joe rocket motors caused the trajectory to be considerably flattened, and as a result the spacecraft was exposed to loading conditions approximately twice those expected for a normal flight.


Spacecraft Control System


The spacecraft control system provides for attitude control and rate stabilization of the spacecraft during the orbital and reentry phases. In addition to the system electronics, the spacecraft control system is composed of two independent reaction control systems (RCS), one of which supplied fuel for the automatic stabilization and control system (ASCS) and fly-by-wire (FBW) modes and the other which, until MA-9, supplied the manual proportional (MP) and the rate stabilization and control system (RSCS) modes. The RSCS unit was installed in the MR-4 and subsequent flights as a backup to one of the secondary modes of the ASCS, that of auxiliary damping. This unit was removed as unnecessary for the MA-9 flight, with major deciding factors being its high fuel- consumption characteristics and weight. The FBW and MP modes were available for direct manual control by the astronaut, initially as backups to the ASCS and in the final two orbital flights as modes of equal priority. Although the control system has operated adequately in all of the manned flights, largely because of the ability of the pilot to exercise precise attitude control manually, this system has exhibited failures of one type or another in nearly every flight. The one exception was the six-pass mission of Astronaut Schirra, in which the system operated correctly.


The single most prevalent malfunction in the control system during the early manned flight program was the intermittent failure of the small 1-pound thrust-chamber assemblies (thrusters). In addition, during a manned suborbital flight (MR-3) a 6-pound thruster also failed to produce thrust when required. During the flight of Astronaut Glenn, intermittent failures of the 1-pound pitch and yaw thrusters would have caused a similar early termination of the mission had the pilot not been present to exercise his manual control option. Immediately following the first inflight thruster failures, a complete analysis was begun to determine the exact cause of the system discrepancy. In the postflight inspections for the MR-3, MA-5, and MA-6 spacecraft, small particles were discovered at critical points in the thrust chamber assembly, and for the MA-5 mission a large metal deposit which partially blocked the thruster orifice was found. Although thruster malfunctions were experienced during the MA-4 flight, the postflight inspection did not reveal any thruster valve contamination. The exact mechanism for transporting these particles, some of which were found to be broken pieces from the stainless-steel dutch-weave screens which distributed the flow, to upstream points is still unknown. Three steps were taken for the MA-7 mission to correct this anomaly, one being the replacement of the dutch-weave screens with a combination of stainless-steel fuel distribution plate and platinum screens, another being the reduction of the bore and size of the heat barrier, and the third being, the relocation of the fuel-metering orifice, to the upstream side of the solenoid valve (ref. 8). While these changes constituted the MA-7 modification, a more refined design change was being developed and qualified in the Project Orbit altitude chamber tests. This configuration, compared in figure 3-10 with previous 1-pound thruster configurations, involved both the 1- and 6- pound thrusters and was installed in the MA-9 spacecraft. No thruster failures of this type occurred on either the MA-7, MA-8, or MA-9 flights after the modifications had been successively incorporated.


55] Figure 3-10. Comparison of 1-pound thrust-chamber configurations.


The horizon scanners, which were used to provide an external reference for the attitude gyros, were a source of difficulty in the earlier orbital flights. In the MR-4 flight after tower jettisoning, the scanner was observed to be generating unexpected ignore signals, the cause of which was later traced to the impingement and heating effects caused by the ignition of the launch-escape rocket. A modification to the horizon-scanner cover eliminated this problem.


In the MA-4 flight, both scanners exhibited output variations which could not be correlated with- attitude changes, and this anomaly was subsequently found to have been partially caused by "cold-cloud effects"; in addition, a shorted capacitor in the scanner circuit contributed to the attitude discrepancy. Since the scanner unit had been designed without accurately taking into account the effect of high-altitude cloud formations in the view field, a temporary modification of altering the bias levels was made for the MA-5 flight, but this change did not completely eliminate the problem. Further system refinement involving signal clipping for the earth portion of the view resulted in a successful modification for the first manned orbital flight. Since that time, only isolated occurrences of "cold-cloud effects" have been observed. During the MA-7 flight, a horizon-scanner circuit failure (see ref. 8) of another type occurred, but because the antenna canister was normally jettisoned prior to landing, it was impossible to conduct a postflight inspection of the hardware and determine the cause of the failure. This malfunction, which occurred in the pitch scanner, is believed to have been random in nature within the scanner circuitry.


The only remaining control system problem of any consequence during the full-scale flight program was the existence of an open circuit in the pitch-rate gyro input to the amplifier-calibrator (Amp-Cal), or autopilot, during the MA-4 mission. The Amp-Cal is the electronic unit which generates automatic control system logic for the various ASCS operating modes. The partial loss of gyro information to the autopilot caused the spacecraft attitude to be in error at retrofire, which in turn resulted in the MA-4 spacecraft's landing some 75 nautical miles up range of the intended point. This malfunction was either not detected during preflight tests or it occurred during the flight.


Although the control system performed satisfactorily during Astronaut Cooper's mission, an electrical short circuit, which occurred at two of the power-carrying plugs into the autopilot and resulted in the loss of the automatic control mode during the final few orbital passes. However, because this malfunction occurred at this specific interface and is primarily of an electrical nature, it is discussed in a later paragraph under Electrical and Sequential Systems. Because of the loss of the automatic control mode during the retrofire and reentry flight maneuvers, the astronaut conducted these maneuvers by using both manual modes available to him.


The only other major modifications to the control system for the 1-day mission of Astronaut Cooper were the addition of a 15-pound capacity fuel tank, which is shown in figure 3-11, and the incorporation of the interconnect valve between the two RCS systems for better fuel utilization, in an emergency, and for more effective fuel jettisoning.


Communications Systems


The original design configuration of the communications systems proved to have been the most conservative of all of the major systems. These systems-the voice transceivers, the radar beacons, the location aids, and the command receivers operated satisfactorily throughout the [56] flight program. Because of the excellent performance of these systems, some of their backup units were deleted, including one of the two command receivers and decoders and the high frequency (HF) recovery transceiver for the MA-8 and MA-9 flights and the ultra-high frequency (UHF) backup voice transceiver for the MA-9 flight. One of the two UHF telemetry transmitters, which were part of the instrumentation system, was also deleted as unnecessary for the MA-9 mission. A slow-scan television system, shown in figure 3-12, was included for evaluation aboard the Faith 7 spacecraft, but the quality and usefulness of its transmissions were not satisfactory.


In the initial two manned orbital flights, it was noted that signals were not being received from the HF recovery transmitter, but because of the circumstances at the time of recovery and the uncertainty of HF reception in the landing area, it could not be established that an anomaly existed. However, when this discrepancy still existed on the MA-8 mission,


Figure 3-11. Auxiliary reaction control system fuel tank.


attention was directed to the ineffectiveness of the HF recovery beacon. Careful analysis revealed that when the HF "whip" antenna was pyrotechnically deployed upon landing, the spacecraft was usually not completely erect in the water. The combination of the electrically conducting products of combustion from the explosive charge used to extend this antenna and the fact that it was extended under water are believed to be the cause of this communications anomaly. The antenna was subsequently deployed by using pressurized nitrogen gas, which is nonconductive, and it was programed such that deployment would not occur until the antenna was clear of the water. Reception from this beacon was satisfactory during the MA-9 mission.


For the MA-8 flight, a pair of more sensitive microphones was installed in the pilot's helmet and the increased sensitivity apparently caused the background noise from the launch vehicle to trigger the voice-operated relay in the air ground circuit. For the MA-9 mission, these microphones were modified to reduce background noise sensitivity such that this triggering action ceased.


Figure 3-12. Television system evaluated during MA-9.


[57] Reports of reception of HF voice communications during the first three manned orbital flight were somewhat inconsistent with regard to quality, but the periods allowed for a complete inflight test of the HF voice equipment were also very brief. At any rate, because of reports that reception of HF voice signals during the first two maimed orbital flights was unsatisfactory' a special HF antenna was installed on the retropackage for the MA-8 flight (see ref. 10). There were reports of excellent reception of signals from this antenna during the flight at ranges exceeding 2,000 nautical miles, while other reports stated that even when the spacecraft was nearly overhead, the reception was poor to unreadable. This inconsistency is not clearly understood, but the effects of spacecraft attitude at the time of transmission, the atmospheric propagation characteristics at the time of contact, and the status of operational ground equipment remain as unknown variables. A more closely controlled test of this special dipole antenna was conducted during the MA-9 flight, and it was fully successful. Although HF voice transmissions were heard during MA-8, the results of MA-9 were more consistent and indicated reliable operation. It might be mentioned that both the


pilots and ground-control personnel preferred the UHF voice equipment to the HF system, particularly since none of the missions were such hat nearly continuous communications were required. The UHF communications, of course, are limited to essentially line-of-sight ranges, but have signal-to-noise characteristics superior to those of HF in flight. However, the MA-9 astronaut found HF communications quite useful during the long periods in which he could not make UHF contact with a network station. Although the command system has never been exercised for a commanded abort, its performance has been entirely satisfactory during other inflight exercises, such as the reception of signals for instrumentation calibration in all orbital flights and for an emergency voice communications test and a commanded wake-up tone in the MA-9 mission. For the unmanned orbital flights, MA-4 and MA-5, the command system was successfully used to control the operation of the spacecraft and bring it safely back from orbit.


Instrumentation System


The instrumentation system monitored over 100 performance variables and events throughout the spacecraft, and the operation of this system was satisfactory throughout the entire Mercury program. The system was designed with enough flexibility to incorporate required instrumentation changes as the program progressed. In the manned orbital flight phase, it was desired to have a more complete temperature survey at discrete spacecraft points, primarily on the spacecraft afterbody; and a low-level commutator circuit was installed. This unit was deleted from the MA-9 spacecraft as having served its purpose and to save weight. The confidence in the telemetry transmitters through the third manned orbital flight led to a decision to eliminate one of the two redundant units from the Faith 7 spacecraft to save weight. The onboard recording capacity for the MA-9 flight was extended by changing the tape speed from 1 7/8 inches per second (ips) to 15/16 ips and reprograming the operation periods such that only essential information was recorded during the expected 34-hour period.


Probably the most widely known system malfunction in the entire Mercury program is that associated with the failure of a limit switch which sensed heat-shield release. During the MA-6 mission, ground- control personnel received a telemetry signal which indicated that the heat shield had been prematurely unlatched from the spacecraft. Although it was believed that this signal was improper and involved an instrumentation failure, a decision was made to reenter with the retropackage attached to insure that the heat shield would not part from the spacecraft during the critical reentry heating period. A postflight examination of the instrumentation revealed that a limit switch had a bent and loose shaft (shown in fig. 3-13) and that manipulation of the sensor without appreciably displacing the sensing shaft would generate an erroneous signal. This experience prompted a change in the installation technique and a directive for tighter quality-control standards to insure that prescribed manufacturing tolerances would be maintained. This type of malfunction did not recur in subsequent flights.


close up of limit switch

[58] Figure 3-13. MA-6 limit switch used to sense heat-shield release.


Early in the flight program, beginning with the Little Joe 5 mission, the mechanical spacecraft clock was found to be sensitive to accelerations in excess of 5g. An electronic digital clock was substituted for this unit and operated satisfactorily.


During the MA-7 mission, the blood-pressure measuring system (BPMS) yielded data which were of only marginal value. The system was thoroughly checked out following the flight, and no major system malfunction was found. It was shown, however, that proper techniques, including establishing a proper amplifier gain setting, correlation with clinically measured values, and the fitting of the pressure cuff to the individual flight astronaut, were not well understood. A thorough review of the entire system, its operating characteristics, and the preflight calibration procedures was conducted in the months after the MA-7 flight, and the data quality for the MA-8 and MA-D missions was correspondingly improved and resulted in usable values. A discussion of this anomaly from a medical standpoint is presented in the Aeromedical Preparations paper.


During the MA-D mission, the programer, which automatically controls the operation and sequence of events of certain spacecraft systems, exhibited two anomalies, one inherent and the other resulting from a structural failure. The inherent anomaly, evident to varying degrees in previous flights, involved a sensitive control circuit containing transistors which actuated power relays to operate the programer. This circuit was sensitive to certain input voltage transients which occasionally caused undesired programer operation. Prior to the MA-9 flight, a loading resistor had been added to reduce the inherent sensitivity, and an on-off switch had been incorporated so that the pilot could shut the system down if improper operation occurred. On two occasions, the unit was inadvertently triggered and continued to call for instrumentation calibrations, one of its programed functions. On both occasions, the [59] astronaut turned the system off, and no serious consequences resulted, but the need to improve system design for future programs in this area, particularly for transistorized circuits, is exemplified.


The other programer anomaly, although in a separate section of the system, involved the shearing of a pin used to maintain alinement of a gear in the programer drive mechanism. Figure 3-14 depicts the misalined gear, which resulted in an inflight binding of the programer and the preclusion of a significant portion of recorded data during the midpoint of the MA-9 flight until the astronaut switched from programed to continuous operation.


During the MA-9 flight, the respiration rate sensor failed to yield reliable data during and after the fifth orbital pass, but other sources of this information were found to be adequate.


close up photo of misaligned gear

Figure 3-14. Misaligned gear in MA-9 programer.


A postflight investigation of the system disclosed a broken solder joint at the attachment point of the sensor lead.


Life-Support Systems


The life-support systems primarily provide for control of the cabin and suit atmospheres, management of metabolic-waste products, and the supply of food and liquid for the astronaut. The major changes to the MA-9 life-support systems, including the environmental control system (ECS) (fig. 3-15), from those of previous missions were accomplished primarily in support of the increased mission time, and the most significant modifications were as follows:


(1) Addition of about 4 pounds of primary breathing oxygen (02), stored under pressure, for a nominal total of 12 pounds in the system.

(2) Increase in the carbon-dioxide (C02) absorber, lithium hydroxide (LiOH), quantity from 4.6 to 5.4 pounds. The amount of activated charcoal, as the odor absorber, was decreased from 1.0 to 0.2 pound, which was sufficient.

(3) Increase in the stored coolant-system water from 39 pounds to 48 pounds.

(4) Increase in the capability of the urine collection and storage system.

(5) Addition of an improved condensate collection and storage system, including a new wick-type condensate trap (shown in fig. 3-16) to extract free water from the suit circuit of the ECS.

(6) Increase of the stored drinking water by 4.5 pounds for a total of 10 pounds of potable water.


Figure 3-15. Environmental control system schematic diagram.

60] Figure 3-16. MA-9 inline condensate trap.

(a) Condensate trap details.

close up of condensation trap

(b) Condensate trap.

condesation trap installed in astronaut's oxygen line

(c) Condensate trap installation.


[61] A parallel coolant control valve (CCV) shown in the upper right corner of figure 3-17 was added in the suit cooling-water circuit for redundancy with the primary valve (top-left on the control plate) in the event of a serious valve blockage by contamination, which was experienced in the MA-8 mission.


The operation of the life-support equipment during the MA-8 mission was normal, except that the suit-circuit CCV was partially blocked by solidified lubricant and delayed the astronaut's stabilization of the cooling system at a comfortable level. Preflight procedures were changed for the MA-9 mission so that the CCV's were cleaned and properly lubricated prior to flight, but after the manned systems tests. The cooling water was also passed through a 0.15 micron filter before being transferred into the spacecraft. Blocking of the CCV during the MA-9 flight was not experienced. However, the astronaut was required to make a large number of minor changes to the suit CCV setting in an attempt to maintain the heat-exchanger dome temperature, which was the cooling system control parameter, within the desired range. No system deficiencies or hardware malfunctions were found during the postflight inspection or testing. It is a characteristic of


Figure 3-17. Redundant coolant valve for MA-9.


the system that changes in metabolic and external suit-circuit heat loads as a result of changes in the astronaut's level of activity, open visor operation, solar heat on the spacecraft, and internal spacecraft equipment heating will be experienced and will be reflected in the coolant requirements for the suit heat exchanger. These heat-load changes are not radical under normal conditions and the corresponding coolant flow changes would be small compared with the capacity of the CCV. It is quite possible that the sensitivity of this small-orifice valve, together with the astronaut s normally varying metabolic heat loads, could have resulted in the need for frequent coolant-flow adjustment.


An inline condensate trap, shown in figure 3-16 was designed to remove excess water from the suit-inlet hose and was installed near the entrance point on the suit. The condensate trap was activated periodically according to the flight plan by the astronaut's opening a hose clamp on the water outlet line from the trap. Condensate water was observed by the astronaut to have been flowing through this line indicating that free water had probably passed around the sponge.


During the 21st orbital pass, the carbon dioxide (CO2) level at the LiOH canister outlet began to show an increase on the C02 meter. Postflight chemical analysis of the canister showed definite channeling of the flow through the canister. Channeling is the localized or restricted passage of gas through the canister, rather than a uniform flow for maximum C02 adsorption. This channeling, which could reduce the effective canister lifetime, has never been experienced during ground testing or during any previous Mercury flight. Based on the amount of unused LiOH at the end of the flight, approximately 27 hours of normal usage remained. However, the actual operating capability of the canister could not be established because of the channeling effects. The exact reason for its occurring on MA-9 could not be established.


The cabin coolant water and fan were turned off according to the flight plan during much of the MA-9 mission in order to evaluate the effectiveness of the cabin cooling circuit. During [62] this time, the electrical load varied according to mission requirements, and the cabin temperature was observed to cycle between 85° F and 95° F, as indicated in figure 3-18. Reduction in the electrical load during this no-cooling period resulted in corresponding reduction in cabin temperature. It is concluded that cabin cooling was not required during periods in which the Mercury spacecraft electrical system was powered down.


Problems were encountered during MA-9 with the condensate transfer system. The needle of the hand-operated pump, used to transfer liquid from the condensate tank to another container, became clogged with metal shavings from the pump shaft and the condensate could not be transferred. Normally, free water removed by the condensate trap and sponge separator flowed directly to the condensate tank, from which it was then intended to be pumped to storage bags. The condensate tank contained a porous plus-to relieve the gas pumped from the sponge into the tank by the action of the sponge separator. Since it was known that this plug could pass water when the tank became nearly filled, the astronaut elected to discontinue operation of the condensate trap when the transfer pump became clogged. This action was taken to stop further flow from the trap to the tank and thereby help to preclude water from being released into the cabin.


No malfunction of the life-support system which compromised the mission or presented a marginal condition to the man occurred during any of the manned Mercury missions. Although minor malfunctions of equipment occurred on


Figure 3-18. Time history of MA-9 spacecraft cabin temperature.


these flights, some of which were alleviated by the astronaut, none of these were repeated on successive flights. The suit cooling system has exhibited a history of undesirable operation, characterized by elevated suit inlet temperatures, wet undergarments, and a general lack of astronaut comfort. However, metabolic heat loads were removed sufficiently to keep body temperatures well below a physiologically marginal value. The causes of these cooling system problems for the suit circuit were twofold:

(1) Selection of an improper cooling system control parameter during the initial design period.

(2) Ineffectiveness of the suit-cooling-circuit water separator because of the unpredicted behavior of free liquid in a weightless condition


Ground testing showed that the steam exhaust duct temperature used in MA-6 and MA-7 missions was not an adequate control parameter for controlling the operation of the heat exchanger. A probe, which sensed the steam temperature at the heat-exchanger dome (see fig. 3-19) between the two coolant evaporating passes, provided a more rapidly responding indication of the heat-exchanger operation. This control temperature parameter was used during the MA-8 and MA-9 flights with satisfactory results. The suit-inlet temperature range of 60° F to 70° F during most of these two flights was more comfortable than the 75°F to 80° F range experienced during MA-6 and MA-7: See figure 3-20 for a summary of suit- inlet temperatures experienced during the four manned orbital flights.


Figure 3-19. Temperature monitoring points on heat exchangers.

63] Figure 3-20. Time history for suit-inlet temperature for manned orbital flights.


Other ground tests showed that water in the suit circuit, when condensed from the gas stream in the heat exchanger, was not carried by the gas flow to the sponge separator. This water is believed to have been held under weightlessness to the metal surfaces by surface tension and flowed from the cooling surfaces to the duct walls, thereby probably passing around the sponge in the separator. The condensate trap, which was installed in the MA-9 ECS, verified the need for a trap which will remove free condensate water traveling along the duct walls. Missions of even longer durations will require the extraction of all free condensate to keep the astronaut's body dry and thereby to obtain maximum comfort and hygiene.


Electrical and Sequential Systems


Except for some early development problems in the sequential system, this system group has performed satisfactorily throughout the Mercury program. Although there were no serious sequential problems throughout the manned flight program, there was an early deployment


of the main parachute during the MR-4 mission and of the drogue parachute during MA-6. The reasons for these premature deployments have never been fully understood, since no system malfunction could be found during exhaustive postflight testing. During the later manned orbital missions, a modification to the sensing circuits for these sequential functions guarded against premature automatic deployment. The contractor was instructed to conduct a single-point failure analysis, which involved a detailed study of the electrical and sequential circuitry to establish all possible failure modes, and this analysis was conducted for all spacecraft systems before the MA-7 flight. The results of this study were evaluated for failure conditions that would singularly jeopardize flight safety, and appropriate modifications were incorporated into the MA-7 and subsequent spacecraft to improve reliability. The greater portion of these changes involved the electrical and sequential systems because of their unique relationship to critical mission functions These changes dictated paralleling of redundant sensing [64] elements in some cases in which the actuation of either element could initiate the proper function. In other cases where it was important that an event signal not be sent early, some elements were changed to a series function, as was done for the parachute-deployment circuitry.


The primary change to the electrical system for the MA-9 mission was the replacement of two 1,500-watt-hour batteries with two 3,000-watt- hour batteries. This change brought the power supply up to one 1,500-watt-hour and five 3,000-watt-hour batteries.


During the early phases of the flight program, difficulty was experienced in maintaining the temperatures of the electrical inverters below the maximum recommended operating level. A cooling system was subsequently installed for the two main inverters, but contamination problems and the limited effectiveness of this cooling system did not alleviate the elevated temperature situation appreciably. However, continued operation of these inverters from mission to mission, in conjunction with ground test results, without experiencing a temperature- associated failure, provided sufficient confidence that these units would operate satisfactorily. Finally, for the MA-9 mission, modified inverters with improved thermal characteristics were installed in place of two of the old style units (main 250 v-amp and 150 v-amp) and the open-cycle evaporative cooling system was deleted. The three spacecraft inverters functioned satisfactorily until late in the MA-9 flight when an electrical short circuit prevented their operating properly.


In the MA-9 flight, the failure which caused the greatest concern was first recognized at the early illumination of the 0.05g sequence light, which indicated that the automatic stabilization and control system (ASCS) had possibly switched to its reentry mode of operation, which would have included the initiation of rate damping and a steady spacecraft roll rate. Subsequent checks by the astronaut revealed, in fact, that this control mode had been enabled: A requirement for a manual retrofire maneuver was therefore imposed on the astronaut, but it was still the plan to use the autopilot during reentry. However, soon after this occurrence the main inverter ceased to supply a-c power and, in the switchover to the standby unit, this redundant element did not start properly. (Refer to fig. 3-21 for details involving the ASCS and power supplies.) Without a-c power for the control system, even the reentry control configuration was disabled; therefore, the astronaut was required to conduct this maneuver with manual control. This task was further complicated by a corresponding loss of gyro attitude indications because of the a-c power failure. A postflight inspection and analysis


Figure 3-21. Relationship of electrical power to control system autopilot.


[65] of the trouble areas disclosed that n short circuit had occurred, both on the power plug (shown in fig. 3-22) to the ASCS amplifier-calibrator and to another connector ( see fig. 3-23), also part of the ASCS power circuit. Both inverters under question were tested thoroughly after the flight and found to operate within specification, indicating that they did not contribute to the malfunction. Strong evidence exists that free water in the spacecraft cabin had been present near the multipin power-plug connection and eventually provided a current path in the insulation between the d-c power and grounding pins shown in right-hand photograph in figure 3-22. Pin N, labeled in the figure, was found to have been completely burned off. Figure 3-23 clearly indicates the significant corrosion revealed on the second connector during the post-flight disassembly and inspection.


Postflight tests duplicated the above hypothesis; that is, a short to ground could be effected upon application of condensate water. Resistance measurements taken across certain pins of the second plug immediately following the flight indicated electrical paths that could have caused the 0.05g indication. A likely source of the liquid which might have caused the electrical short circuit was the porous vent of the condensate tank in the environmental control system.


This tank is located in the proximity of the autopilot power plugs, and normal cycling of the sponge squeezer during the flight could have forced condensate through the vent. Another possible source of water which could have produced the short circuit is the local condensation of cabin humidity, which may have been present because of a leak in the drinking- water valve or because of water vapor exhaled by the pilot when his helmet faceplate was open. Or the water droplets which leaked from the valve may have somehow been deposited, in part, directly on the power plug. This experience points up the need to minimize or eliminate the presence of free liquid or high humidity in a spacecraft cabin where electrical systems are functioning and to insulate and seal bare electrical connectors more effectively.


Figure 3-22. Postflight photograph of MA-9 auto-pilot powerplug.

(a) Front view showing burnt pin.

close up of cracked connector plug

(b) Rear view showing x-rayed current paths in insulation.


close up of electrical connector

Figure 3-23. Postflight photograph of MA-9 connector-socket rear face.

[66] Concluding Remarks


The Mercury spacecraft systems design and development phases were conducted concurrently and although this philosophy involved a known risk, it made possible the early realization of the project objectives. During this time, many valuable lessons were learned and exploited in the development and operation of manned space-flight systems.


In the system design, maximum use was made of existing technology and off-the-shelf equipment, and systems concepts were kept simple. However, some important advances in the technology also had to be initiated. It was found that the spacecraft and its systems must be designed for operational conditions. Examples of the design-for-operation standard relating to the preflight activities are system accessibility and the simplification of system interfaces. It is also important in the early system design to allow for an inevitable growth in weight.


During development and qualification testing, the test criteria cannot be compromised in most instances, since an overlooked system inefficiency will inevitably show up later where a redesign is more costly. However, it was also found in Mercury that no single qualification criterion necessarily applies to all systems, and local operational conditions must be individually evaluated for each system. Whenever system components are significantly modified, as was done for the Faith 7 spacecraft to make possible the 34-hour flight capability, a new ground test program for hardware requalification should be administered to insure maintenance of previous reliability and operational standards.


In the area of hardware operation and performance evaluation, the Mercury flight program has been a most valuable experience. The most important lesson learned from operation of the spacecraft control system is that the pilot is a reliable backup to automatic system modes. In fact, the pilot's ability to control accurately the spacecraft attitude was instrumental in three of the four manned orbital flights in completing the mission successfully when a malfunction was present in the automatic system. Another valuable lesson in both the control system and cooling system designs was the avoidance of components which are especially sensitive to contamination. The small valves used to meter reaction control fuel and environmental control system cooling water should have been designed to employ larger flow areas to reduce susceptibility to particle blockage. Other than guarding against stray voltages and sensitivity to transients, the major lesson derived from the performance of the electrical . and sequential systems was the need: to seal and insulate effectively all electrical connectors from possible sources of free liquid and humidity in the spacecraft cabin. In the life support system, it was also found that the cooling systems must be designed with adequate margins and that food, water, and waste management devices require particular attention because of plumbing complexity and the effects: of weightlessness.


Throughout the Mercury development and flight programs, quality control and rigid manufacturing standards were found to be absolutely mandatory if incidental flight failure and discrepancies were to be avoided. Throughout the project, a careful and continuing attention was given to engineering detail in order to make possible the early recognition of system weaknesses and their implications in the operation of flight hardware and to provide meaningful and effective courses of action. This attention to detail was an important reason for the success of the Mercury flight program, particularly the manned suborbital and orbital missions.


Acknowledgement. The authors wish to gratefully acknowledge the analytical and documentary efforts of the many NASA engineers and technicians who applied their knowledge and foresight unselfishly during the postflight evaluations of the various spacecraft systems for each Mercury mission and without whose contributions this paper would not have been possible.


[67] References


1. FAGET M. A., and PILAND, R. O.: Mercury Capsule and its Flight Systems. IAS Paper No. 60-34, Presented at IAS 28th Annual Meeting (New York, N.Y.), Jan. 25-27, 1960.

2. ANDERTON, DAVID A.: How Mercury Capsule Design Evolved. Aviation Week, Vol. 74, No. 21, May 22, 1961.

3. BOND, ALECK C.: Mercury Spacecraft Systems. Proc. Conf. on Results of the First U.S. Manned Suborbital Space Flight. NASA, Nat. Inst. Health, Nat. Acad. Sci., June 6, 1961, pp. 11-18.

4. HAMMACK JEROME B.: Spacecraft and Flight Plan for Mercury-Redstone 4 Flight. Results of the Second U.S. Manned Suborbital Space Flight, July 21, 1961. Supt. Doc., U.S. Government Printing Office (Washington, D.C.), pp. 3 S.

5. KLEINKNECHT, KENNETH S., BLAND, WILLIAM M., Jr., and FIELDS, D. M.: Spacecraft and Spacecraft Systems. Results of the First U.S. Manned Orbital Space Flight, February 20, 1962. Supt. Doc., U.S. Government Printing Office (Washington, D.C.), pp. 5-30.

6. JOHNSTON, RICHARD S., SAMONSKI, FRANK H., Jr., LIPPITT, MAXWELL NV., and RADNOFSKY, MATTHEW I.: Life Support Systems and Biomedical Instrumentation. Results of the First U.S. Manned Orbital Space Flight, February 20, 1962. Supt Doc., U.S. Government Printing Office (Washington, D.C.), pp. 31-44.

7. VOAS, ROBERT B.: Manual Control of the Mercury Spacecraft. Astronautics, Vol. 7, No. 3, March 1962, pp. 18-20 and 34-38.

8. BOYNTON, JOHN H., and FIELDS, E. M.: Spacecraft and Launch-Vehicle Performance. Results of the Second United States Manned Orbital Space Flight, May 24, 1962. NASA SP-6, Supt. Doc., U.S. Government Printing Office (Washington, D.C.), pp. 1-14.

9 ANON: Mercury Provides Data for Apollo, Gemini. Aviation Week, Vol. 77, No. 1, July 2, 1962, pp. 98 99.

10. BOYNTON, JOHN H., and FISHER, LEWIS R.: Spacecraft and Launch-Vehicle Performance. Results of the Third United States Manned Orbital Space Flight, October 3, 1962. NASA SP-12, Supt. Doc., U.S. Government Printing Office ( Washington, D.C. ), pp. 1-11.

11. BLAND, WILLIAM M., JR., and BERRY, CHARLES A.: Project Mercury Experiences. Astronautics and Aerospace Engineering, Vol. 1, No. 1, Feb. 1963, pp. 29-34.

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