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The purpose of the Committee was to take a long-term look at man-in-space problems, leading eventually to recommendations on future missions and on broad aspects of Center research programs to ensure that the Centers were providing proper information. Committee investigations would range beyond Mercury and Dyna-Soar but would not be overly concerned with specific vehicular configurations. The Committee would report directly to the Office of Aeronautical and Space Research.
Bruce T. Lundin of the Lewis Research Center reported to members on propulsion requirements for various modes of manned lunar landing missions, assuming a 10,000-pound spacecraft to be returned to earth. Lewis mission studies had shown that a launch into lunar orbit would require less energy than a direct approach and would be more desirable for guidance, landing reliability, etc. From a 500,000 foot orbit around the moon, the spacecraft would descend in free fall, applying a constant-thrust decelerating impulse at the last moment before landing. Research would be needed to develop the variable-thrust rocket engine to be used in the descent. With the use of liquid hydrogen, the launch weight of the lunar rocket and spacecraft would be 10 to 11 million pounds. Additional Details: here....
In addition, Goett informed the Committee that the Vega had been eliminated as a possible booster for use in one of the intermediate steps leading to the lunar mission. The primary possibility for the earth satellite mission was now the first-generation Saturn and for the lunar flight the second-generation Saturn.
On February 19, NASA officials again presented the ten-year timetable to the House Committee. A lunar soft landing with a mobile vehicle had been added for 1965. On March 28, NASA Administrator T. Keith Glennan described the plan to the Senate Committee on Aeronautical and Space Sciences. He estimated the cost of the program to be more than $1 billion in Fiscal Year 1962 and at least $1.5 billion annually over the next five years, for a total cost of $12 to $15 billion. Additional Details: here....
To open these discussions, Director Robert R. Gilruth summarized the guidelines: manned lunar reconnaissance with a lunar mission module, corollary earth orbital missions with a lunar mission module and with a space laboratory, compatibility with the Saturn C-1 or C-2 boosters (weight not to exceed 15,000 pounds for a complete lunar spacecraft and 25,000 pounds for an earth orbiting spacecraft), 14-day flight time, safe recovery from aborts, ground and water landing and avoidance of local hazards, point (ten square-mile) landing, 72-hour postlanding survival period, auxiliary propulsion for maneuvering in space, a "shirtsleeve" environment, a three-man crew, radiation protection, primary command of mission on board, and expanded communications and tracking facilities. In addition, a tentative time schedule was included, projecting multiman earth orbit qualification flights beginning near the end of the first quarter of calendar year 1966.
NASA Deputy Administrator Hugh L. Dryden announced that the advanced manned space flight program had been named "Apollo." George M. Low, NASA Chief of Manned Space Flight, stated that circumlunar flight and earth orbit missions would be carried out before 1970. This program would lead eventually to a manned lunar landing and a permanent manned space station. Additional Details: here....
"In order to prepare such a program, I have formed a small working group, consisting of Eldon Hall, Oran Nicks, John Disher, and myself. This group will endeavor to establish ground rules for manned lunar landing missions; to determine reasonable spacecraft weights; to specify launch vehicle requirements; and to prepare an integrated development plan, including the spacecraft, lunar landing and takeoff system, and launch vehicles. This plan should include a time-phasing and funding picture, and should identify areas requiring early studies by field organizations."
Fundamental decisions were made as a result of this and a previous meeting on September 20.. Additional Details: here....
At an executive session, Air Force and NASA programs of orbital rendezvous, refueling, and descent from orbit were discussed. Long-range Air Force studies on a lunar base were in progress as well as research on more immediate missions, such as rendezvous by an unmanned satellite interceptor for inspection purposes, manned maintenance satellites, and reentry methods. NASA plans for the manned lunar landing mission included the possible use of the Saturn booster in an orbital staging operation employing orbital refueling. Reentry studies beyond Mercury were concentrated on reentry at escape speeds and on a spacecraft configuration capable of aerodynamic maneuvering during reentry.
(Heating), Roger A. Anderson (Structures and Materials), Wilford E. Sivertson, Jr. (Instrumentation and Communications), David Adamson (Human Factors), and Joseph G. Thibodaux, Jr. (Onboard Propulsion).
The Group for Instrumentation and Communications discussed a set of working guidelines on spacecraft instrumentation and communications, tracking considerations, and deep-space communication requirements. Progress of the three Apollo feasibility study contracts was reviewed and the proposed MIT Lincoln Laboratory study on a systems concept for the ground instrumentation and tracking required for the Apollo mission was discussed. Reports of studies were given by members from the NASA Centers. The Group recommendations were :
After reviewing the status of the contractors' Apollo feasibility studies, the Group on Trajectory Analysis discussed studies being made at NASA Centers. An urgent requirement was identified for a standard model of the Van Allen radiation belt which could be used in all trajectory analysis related to the Apollo program,
The Group on Heating, after consideration of NASA and contractor studies currently in progress, recommended experimental investigation of control surface heating and determination of the relative importance of the unknowns in the heating area by relating estimated "ignorance" factors to resulting weight penalties in the spacecraft. The next day, three members of this Group met for further discussions and two areas were identified for more study: radiant heat inputs and their effect on the ablation heatshield, and methods of predicting heating on control surfaces, possibly by wind tunnel tests at high Mach numbers.
The Group on Human Factors considered contractors' studies and investigations being done at NASA Centers. In particular, the Group discussed the STG document, "Project Apollo Life Support Programs," which proposed 41 research projects. These projects were to be carried out by various organizations, including NASA, DOD, industry, and universities. Medical support experience which might be applicable to Apollo was also reviewed.
The Group on Structures and Materials, after reviewing contractors' progress on the Apollo feasibility studies, considered reports on Apollo-related activities at NASA Centers. Among these activities were work on the radiative properties of material suitable for temperature control of spacecraft (Ames), investigation of low-level cooling systems in the reentry module (Langley), experiments on the landing impact of proposed reentry module shapes (Langley), meteoroid damage studies (Lewis), and the definition of suitable design criteria and safety factors to ensure the structural integrity of the spacecraft STG.
The Group on Configurations and Aerodynamics recommended :
The engineering sketch drawn by John D. Bird of Langley Research Center on May 3, 1961, indicated the thinking of that period: By launching two Saturn C-2's, the lunar landing mission could be accomplished by using both earth rendezvous and lunar rendezvous at various stages of the mission.
The evaluation would consider:
At the same time, Robert R. Gilruth was named Director of the Manned Spacecraft Center to be located in Houston, Tex. The Directors of NASA's nine field centers would, like the newly appointed program Directors, report to Seamans.
The Sub-Board would :
Four Saturn-Apollo Coordination Panels were established to make available the technical competence of MSFC and STG for the solution of interrelated problems of the launch vehicle and the spacecraft. The four included the Launch Operations, Mechanical Design, Electrical and Electronics Design, and Flight Mechanics, Dynamics, and Control Coordination Panels. Although these Panels were designated as new Panels, the members selected by STG and MSFC represented key technical personnel who had been included in the Mercury-Redstone Panels, the Mercury-Atlas Program Panels, the Apollo Technical Liaison Groups, and the Saturn working groups. The Charter was signed by von Braun and Gilruth. Charter of the MSFC-STG Space Vehicle Board, October 3, 1961.
In further discussion, Paul J. DeFries of Marshall Space Flight Center MSFC presented a list of proposed guidelines for use in studying early manned lunar landing missions:
David G. Hoag, MIT, personal notes, October 1961..
In addition, the Apollo Project Office, which had been part of the MSC Flight Systems Division, would now report directly to the MSC Director and would be responsible for planning and directing all activities associated with the completion of the Apollo spacecraft project. Primary functions to be performed by the Office would include:
Letter contract No. NAS 9-150, authorizing work on the Apollo development program to begin on January 1, 1962, was signed by NASA and NAA on December 21. Under this contract, NAA was assigned the design and development of the command and service modules, the spacecraft adapter, associated ground support equipment, and spacecraft integration. Formal signing of the contract followed on December 31.
The center couch, including the crewman parachute and survival kit, could be folded out to a sleep position and stowed under either remaining couch. Allowance was made for the crewman to turn over.
Principal problems remaining were the difficulty of removing the center couch and providing the clearances needed for the couch positions specified for various phases of the lunar mission.
Robert O. Piland, Deputy Project Manager
William F. Rector, Special Assistant
Calvin H. Perrine, Flight Technology
Lee N. McMillion, Crew Systems
David L. Winterhalter, Sr., Power Systems
Wallace D. Graves, Mechanical Systems
Milton C. Kingsley, Electrical Systems
(Vacant), Ground Support Equipment
Jack Barnard, Apollo Office at MIT
(Vacant), Reliability and Quality Control
Emory F. Harris, Operations Requirements
Robert P. Smith, Launch Vehicle Integration
Owen G. Morris, Mission Engineering
Marion R. Franklin, Ground Operational Support Systems
Alan B. Kehlet, Engineering
Alan B. Kehlet, Acting Manager, Quality Control and Engineering
Herbert R. Ash, Acting Manager, Business Administration
The launch vehicle required was a single Saturn C-5, consisting of the S-IC, S-II, and S-IVB stages. To provide a maximum launch window, a low earth parking orbit was recommended. For greater reliability, the two-stage-to-orbit technique was recommended rather than requiring reignition of the S-IVB to escape from parking orbit.
The current concepts of the Apollo command and service modules would not be altered. The lunar excursion vehicle (LEV), under intensive study in 1961, would be aft of the service module and in front of the S-IVB stage. For crew safety, an escape tower would be used during launch. Access to the LEV would be provided while the entire vehicle was on the launch pad.
Both Apollo and Saturn guidance and control systems would be operating during the launch phase. The Saturn guidance and control system in the S-IVB would be "primary" for injection into the earth parking orbit and from earth orbit to escape. Provisions for takeover of the Saturn guidance and control system should be provided in the command module. Ground tracking was necessary during launch and establishment of the parking orbit, MSFC and GSFC would study the altitude and type of low earth orbit.
The LEV would be moved in front of the command module "early" in the translunar trajectory. After the S-IVB was staged off the spacecraft following injection into the translunar trajectory, the service module would be used for midcourse corrections. Current plans were for five such corrections. If possible, a symmetric configuration along the vertical center line of the vehicle would be considered for the LEV. Ingress to the LEV from the command module should be possible during the translunar phase. The LEV would have a pressurized cabin capability during the translunar phase. A "hard dock" mechanism was considered, possibly using the support structure needed for the launch escape tower. The mechanism for relocation of the LEV to the top of the command module required further study. Two possibilities were discussed: mechanical linkage and rotating the command module by use of the attitude control system. The S-IVB could be used to stabilize the LEV during this maneuver.
The service module propulsion would be used to decelerate the spacecraft into a lunar orbit. Selection of the altitude and type of lunar orbit needed more study, although a 100-nautical-mile orbit seemed desirable for abort considerations.
The LEV would have a "point" landing (±½ mile) capability. The landing site, selected before liftoff, would previously have been examined by unmanned instrumented spacecraft. It was agreed that the LEV would have redundant guidance and control capability for each phase of the lunar maneuvers. Two types of LEV guidance and control systems were recommended for further analysis. These were an automatic system employing an inertial platform plus radio aids and a manually controlled system which could be used if the automatic system failed or as a primary system.
The service module would provide the prime propulsion for establishing the entire spacecraft in lunar orbit and for escape from the lunar orbit to earth trajectory. The LEV propulsion system was discussed and the general consensus was that this area would require further study. It was agreed that the propulsion system should have a hover capability near the lunar surface but that this requirement also needed more study.
It was recommended that two men be in the LEV, which would descend to the lunar surface, and that both men should be able to leave the LEV at the same time. It was agreed that the LEV should have a pressurized cabin which would have the capability for one week's operation, even though a normal LOR mission would be 24 hours. The question of lunar stay time was discussed and it was agreed that Langley should continue to analyze the situation. Requirements for sterilization procedures were discussed and referred for further study. The time for lunar landing was not resolved.
In the discussion of rendezvous requirements, it was agreed that two systems be studied, one automatic and one providing for a degree of manual capability. A line of sight between the LEV and the orbiting spacecraft should exist before lunar takeoff. A question about hard-docking or soft-docking technique brought up the possibility of keeping the LEV attached to the spacecraft during the transearth phase. This procedure would provide some command module subsystem redundancy.
Direct link communications from earth to the LEV and from earth to the spacecraft, except when it was in the shadow of the moon, was recommended. Voice communications should be provided from the earth to the lunar surface and the possibility of television coverage would be considered.
A number of problems associated with the proposed mission plan were outlined for NASA Center investigation. Work on most of the problems was already under way and the needed information was expected to be compiled in about one month.
(This meeting, like the one held February 13-15, was part of a continuing effort to select the lunar mission mode).
A study on the direct vision requirement for lunar landing showed that, to have a simultaneous direct view of the lunar landing point and the landing feet without changing the spacecraft configuration, a periscope with a large field of view integrated with a side window would be needed. A similar requirement on the general-purpose telescope could thus be eliminated, reducing the complexity of the telescope design.
Another study showed that, with an additional weight penalty of from five to ten pounds, an optical drift indicator for use after parachute deployment could easily be incorporated into the general-purpose telescope.
Concurrently, a number of NAA latching concepts were in preparation for presentation to NASA, including that of an outward-opening, quick- opening crew door without an outer emergency panel. This design, however, had weight and complexity disadvantages, as well as requiring explosive charges.
For the service module RCS, a quadruple arrangement was chosen which was basically similar to the command module RCS except that squib valves and burst discs were eliminated.
In addition, Gilruth noted that a reevaluation of the Saturn C-1 and C-1B launch capabilities appeared to indicate that neither vehicle would be able to test the complete Apollo spacecraft configuration, including the lunar excursion module. Complete spacecraft qualification would require the use of the Saturn C-5.
One 16-mm camera was also planned for the spacecraft. This camera would be positioned level with the commander's head and directed at the main display panel. It could be secured to the telescope for recording motion events in real time such as rendezvous, docking, launch and recovery of a lunar excursion module, and earth landing; it could be hand-held for extravehicular activity.
The command module (CM) would now be required to provide the crew with a one-day habitable environment and a survival environment for one week after touching down on land or water. In case of a landing at sea, the CM should be able to recover from any attitude and float upright with egress hatches free of water. Additional Details: here....
Another closed-hatch configuration under consideration would entirely eliminate the CM airlock. Astronauts transferring to and from the lunar excursion module would be in a pressurized environment constantly.
The mission constraints to be used for this study were :
"Man, in his quest for knowledge and progress, is determined and cannot be deterred. The exploration of space will go ahead, whether we join in it or not, and it is one of the great adventures of all time, and no nation which expects to be the leader of other nations can expect to stay behind in this race for space. . . .
"We choose to go to the moon in this decade and do the other things, not because they are easy, but because they are hard, because that goal will serve to organize and measure the best of our energies and skills, because that challenge is one that we are willing to accept, one we are unwilling to postpone, and one which we intend to win, and the others, too.
"It is for these reasons that I regard the decision last year to shift our efforts in space from low to high gear as among the most important decisions that will be made during my incumbency in the office of the Presidency. . . ."
Among the items deleted from the command module (CM) were exercise and recreation equipment, personal parachutes and parachute containers located in the couches, individual survival kits, solar radiation garments, and eight-ball displays. A telescope, cameras and magazines considered scientific equipment, and a television monitor were deleted from the CM instrumentation system.
The Office of Systems under NASA's Office of Manned Space Flight completed a manned lunar landing mode comparison embodying the most recent studies by contractors and NASA Centers. The report was the outgrowth of the decision announced by NASA on July 11 to continue studies on lunar landing modes while basing planning and procurement primarily on the lunar orbit rendezvous (LOR) technique. Additional Details: here....
Pad aborts: PA-1, PA-2, etc.
Missions using Little Joe II launch vehicles: A-001, A-002, etc. Missions using Saturn C-1 launch vehicles: A-101, A-102, etc. Missions using Saturn C-1B launch vehicles: A-201, A-202, etc. Missions using Saturn C-5 launch vehicles: A-501, A-502, etc.
The 'A' denoted Apollo, the first digit stood for launch vehicle type or series, and the last two digits designated the order of Apollo spacecraft flights within a vehicle series.
A review of body angles used for the current couch geometry disclosed that the thigh-to-torso angle could be closed sufficiently for a brief period during reentry to shorten the overall couch length by the required travel along the Z-Z axis. The more acute angle was desirable for high g conditions. This change in the couch adjustment range, as well as a revision in the lower leg angle to gain structure clearance, would necessitate considerable couch redesign.
The tracking network would consist of stations equipped with 9-meter (30foot) antennas for near-earth tracking and communications and of stations having 26-meter (85-foot) antennas for use at lunar distances. A unified S-band system, capable of receiving and transmitting voice, telemetry, and television on a single radio-frequency band, was the basis of the network operation.
On March 12, 1963, during testimony before a subcommittee of the House Committee on Science and Astronautics, Edmond C. Buckley, Director of OTDA, described additional network facilities that would be required as the Apollo program progressed. Three Deep Space Instrumentation Facilities with 26-meter (85- foot) antennas were planned: Goldstone, Calif. (completed); Canberra, Australia (to be built); and a site in southern Europe (to be selected). Three new tracking ships and special equipment at several existing network stations for earth-orbit checkout of the spacecraft would also be needed.
Four days earlier, MSC had added specifications for an extravehicular suit communications and telemetry (EVSCT) system to the space suit contract with Hamilton Standard Division of United Aircraft Corporation. The EVSCT system included equipment for three major operations:
Also, Collins awarded a contract to the Leach Corporation for the development of command and service module (CSM) data storage equipment. The tape recorders must have a five-hour capacity for collection and storage of data, draw less than 20 watts of power, and be designed for in-flight reel changes.
Also investigated was crew transfer from the CM to the LEM, to determine the requirements for crew performance and, from this, to define human engineering needs. North American concluded that a separate LEM airlock was not needed but that the CSM oxygen supply system's capacity should be increased to effect LEM pressurization.
On November 29, North American presented the results of docking simulations, which showed that the free flight docking mode was feasible and that the 45-kilogram (100-pound) service module (SM) reaction control system engines were adequate for the terminal phase of docking. The simulations also showed that overall performance of the maneuver was improved by providing the astronaut with an attitude display and some form of alignment aid, such as probe.
RCA Service Company to design and build two vacuum chambers at MSC. The facility was used in astronaut training and spacecraft environmental testing. using carbon arc: lamps, the chambers simulated the sun's intensity, permitting observation of the effects of solar heating encountered on a lunar mission. At the end of July, MSC awarded RCA another contract (worth $3,341,750) for these solar simulators.
(Apocynthion and pericynthion are the high and low points, respectively, of an object in orbit around the moon (as, for example, a spacecraft sent from earth). Apolune and perilune also refer to these orbital parameters, but these latter two words apply specifically to an object launched from the moon itself.)
Specialty areas for the second generation were: trainers and simulators, Neil A. Armstrong; boosters, Frank Borman; cockpit layout and systems integration, Charles Conrad, Jr.; recovery system, James A. Lovell, Jr.; guidance and navigation, James A. McDivitt; electrical, sequential, and mission planning, Elliot M. See, Jr.; communications, instrumentation, and range integration, Thomas P. Stafford; flight control systems, Edward H. White II; and environmental control systems, personal equipment, and survival equipment, John W. Young.
Ferrando and Lineberry found that, once abort factors are considered, there exist "very few" orbits that are acceptable from which to begin the descent. They reported that the most advantageous orbit for the CSM would be a 147-kilometer (80-nautical-mile) circular one.
Membership on the panel included representatives from MSC, MSFC, NASA Headquarters, North American, Grumman, and MIT, with other NASA Centers being called on when necessary. By outlining the most accurate mission plan possible, the panel would ensure that the spacecraft could satisfy Apollo's anticipated mission objectives. Most of the panel's influence on spacecraft design would relate to the LEM, which was at an earlier stage of development than the CSM. The panel was not given responsibility for preparing operational plans to be used on actual Apollo missions, however.
Also, Grumman advised that, from the standpoint of landing stability, a five-legged LEM was unsatisfactory. Under investigation were a number of landing gear configurations, including retractable legs.
Other items received considerable attention: A six-foot umbilical hose would be adequate for the astronaut in the CM. The location of spacecraft water, oxygen, and electrical fittings was judged satisfactory, as were the new couch assist handholds. The astronaut's ability to operate the environmental control system (ECS) oxygen flow control valve while couched and pressurized was questionable. Therefore, it was decided that the ECS valve would remain open and that the astronaut would use the suit control valve to regulate the flow. It was also found that the hand controller must be moved about nine inches forward.
Through this simulation, Grumman sought primarily to evaluate the astronauts' ability to perform the landing maneuver manually, using semiautomatic as well as degraded attitude control modes. Other items evaluated included the flight control system parameters, the attitude and thrust controller configurations, the pressure suit's constraint during landing maneuvers, the handling qualities and operation of LEM test article 9 as a freeflight vehicle, and manual abort initiation during the terminal landing maneuver.
Use of the fixed couch required relocation of the main and side display panels and repositioning of the translational and rotational hand controllers. During rendezvous and docking operations, the crew would still have to adjust their normal body position for proper viewing.
There were a number of other cases wherein North American and ASPO agreed on procedures which simply required formal statements of what would be done. Examples of these were:
Recommendations of the board were not binding. If a Center Director decided against a board recommendation, he would, however, discuss and clear the proposed action with the Director of OMSF.
When the Panel Review Board assumed its duties, the Space Vehicle Review Board was abolished.
"The consequences," Bryant concluded, "of imposing an ever-increasing number of these flight restrictions is obvious - the eventual loss of almost all operational flexibility. The only solution is . . . (a) meticulous examination of every constraint which tends to reduce the number of available launch opportunities," looking toward eliminating "as many as possible."
The document called for five flight and three test articles. The Lunar Orbiter spacecraft would be capable of photographing the moon from a distance of 22 miles above the surface. Overall cost of the program was estimated at between $150 and $200 million.
Also Grumman reported that a preliminary analysis showed the reaction control system plume heating of the LEM landing gear was not a severe problem. (This difficulty had been greatly alleviated by the change from five to four landing legs on the vehicle.
On the following day, the Center informed North American also that a new mechanical clock timer system would be provided in the CM for indicating elapsed time from liftoff and predicting time to and duration of various events during the mission.
Recent load analysis at North American placed the power required for a 14-day mission at 577 kilowatt-hours, a decrease of about 80 kilowatt-hours from earlier estimates.
On October 1, North American delivered the test fixture to the U.S. Navy Aviation Medical Acceleration Laboratory, where the first phase of the manned centrifuge program was scheduled to begin that month.
On October 11, MSC Crew Systems Division (CSD) tested the suit's mobility with the portable life support system (PLSS). CSD researchers found that the PLSS did not restrict the wearer's movement because the suit supported the weight of the PLSS. Shifts in the center of gravity appeared insignificant. The PLSS controls, because of their location, were difficult to operate, which demanded further investigation.
At this same meeting, Grumman presented a comparison of radially and laterally folded landing gears (both of 457-centimeter (180-inch) radius). The radial-fold configuration, MSC reported, promised a weight savings of 22-2 kilograms (49 pounds). MSC approved the concept, with an 876-centimeter (345-inch) adapter. Further, an adapter of that length would accommodate a larger, lateral fold gear (508 centimeters (200 inches)), if necessary. During the next several weeks, Grumman studied a variety of gear arrangements (sizes, means of deployment, stability, and even a "bending" gear). At a subsequent LEM Mechanical Systems Meeting, on November 10, Grumman presented data (design, performance, and weight) on several other four-legged gear arrangements - a 457-centimeter (180-inch), radial fold "tripod" gear (i.e., attached to the vehicle by three struts), and 406.4-centimeter (160-inch) and 457-centimeter (180-inch) cantilevered gears. As it turned out, the 406.4-centimeter (160-inch) cantilevered gear, while still meeting requirements demanded in the work statement, in several respects was more stable than the larger tripod gear. In addition to being considerably lighter, the cantilevered design offered several added advantages:
These guidelines had been arrived at after extensive consultation within NASA as a whole as well as with the scientific community.
Mission rules established crew safety as the major consideration in all mission decisions and detailed actions to be taken in the event of a failure in any system or subsystem.
Associate Administrator for Manned Space Flight George E. Mueller had recommended the changeover from the Saturn I to the Saturn IB to NASA Administrator James E. Webb on October 26. Webb's concurrence came two days later.
On November 18, Mueller further defined the flight schedule planning. Early Saturn IB flights might not be able to include the LEM, but every effort must be made to phase the LEM into the picture as early as possible. Launch vehicle payload capability must be reached as quickly as practicable. Subsystems for the early flights should be the same as those intended for lunar missions. To conserve funds, the first Saturn V vehicle would be used to obtain reentry data early in the Saturn test program.
Despite the contractor's findings, MSC concluded that there was no need for an RE warning system aboard the spacecraft, believing that radiation warning could be handled more effectively by ground systems. But MSC did concur in the recommendation for a combined proton direction and external environment detection system and authorized North American to proceed with its design and development.
On this same day, MSC awarded a $183,152 contract to Wyle Laboratories to construct a high-intensity acoustic facility, also for testing spacecraft parts. The facility would generate noise that might be encountered in space flight.
Maxime A. Faget, Assistant Director for Engineering and Development Christopher C. Kraft, Jr., Assistant Director for Flight Operations Donald K. Slayton, Assistant Director for Flight Crew Operations Wesley L. Hjornevik, Assistant Director for Administration Joseph F. Shea, Manager, Apollo Spacecraft Program Office Charles W. Mathews, Manager, Gemini Program Office and G. Merritt Preston, Manager, MSC Florida Operations.
All sequencing was normal. The tower-jettison motor sent the escape tower into a proper ballistic trajectory. The drogue parachute deployed as programmed, followed by the pilot parachute and main parachutes. The test lasted 165.1 seconds. The postflight investigation disclosed only one significant problem: exhaust impingement that resulted in soot deposits on the CM.
MSC directed North American to proceed with the tower flap as its prime effort, and attempt to solve the stability problem at the earliest possible date. MSC's Engineering and Development Directorate resumed its study of both configurations, with an in-depth analysis of the canard system, in case the stability problem on the tower flap could not be solved by the end of the year.
There would be two ways for the astronauts to get from one spacecraft to the other. The primary mode involved docking and passage through the transfer tunnel. An emergency method entailed crew and payload transfer through free space. The CSM would take an active part in translunar docking, but both spacecraft must be able to take the primary role in the lunar orbit docking maneuver. A single crewman must be able to carry out the docking maneuver and crew transfer.
The ASPO Manager's proposal resulted from experience that had arisen because of unfortunate terminology used to designate the extra fuel. Originally the fuel budget for various phases of the mission had been analyzed and a 10 percent allowance had been made to cover - at that time, unspecified - contingencies, dispersions, and uncertainties. Mistakenly this fuel addition became known as a "10% reserve"! John P. Mayer and his men in the Mission Planning and Analysis Division worried because engineers at North American, Grumman, and NASA had "been freely 'eating' off the so-called 'reserve'" before studies had been completed to define what some of the contingencies might be and to apportion some fuel for that specific situation. Mayer wanted the item labeled a "10% uncertainty."
Shea recommended also that the capacity of the LEM descent tanks be sufficient to achieve an equiperiod orbit, should this become desirable. However, the spacecraft should carry only enough propellant for a Hohmann transfer. This was believed adequate, because the ascent engine was available for abort maneuvers if the descent engine failed and because a low altitude pass over the landing site was no longer considered necessary. By restricting lunar landing sites to the area between ±5 degrees latitude and by limiting the lunar stay time to less than 48 hours, a one-half-degree, rather than two-degree, plane change was sufficient.
In the meantime, Shea reported, his office was investigating how much weight could be saved by these propellant reductions.
Penetration of the thicker targets was about 13.970 millimeters (0.55 inch). In the thinner targets, the ablator was pierced. Debris tore through the steel honeycomb and produced pinholes on the rear steel sheet. Damage to the ablator was confined to two or three honeycomb cells and there was no cracking or spalling on the surface.
Tests at Ames of thermal performance of the ablation material under high shear stress yielded favorable preliminary results.
The selection criteria at all stages were determined by lunar surface requirements prepared by OMSF. Fryklund emphasized that a landing at the least hazardous spot, rather than in the area with the most scientific interest, was the chief aim of the site selection process.
On this same date, Grumman concluded negotiations with Allison Division of General Motors Corporation for design and fabrication of the LEM descent engine propellant storage tanks (at a cost of $5,479,560).
Resizing of the LEM propulsion tanks was completed by Grumman. The cylindrical section of the descent tank was extended 34.04 millimeters (1.34 inches), for a total of 36.27 centimeters (14.28 inches) between the spherical end bells. The ascent tanks (two-tank series) were 1240.54 centimeters (48.84 inches) in diameter.
Intermediate objectives for the Apollo program were outlined: the qualification of a manned CSM capable of earth reentry at parabolic velocities after an extended space mission; qualification of a manned LEM both physically and functionally compatible with the CSM; and demonstration of manned operations in deep space, including lunar orbit. The most significant basic test plan objective formulated during the study was the need for flexibility to capitalize on unusual success or to compensate for unexpected difficulties with minimum impact on the program.
Only one major issue in the test plan remained unresolved - lunar descent radar performance and actual lunar touchdown. Two possible solutions were suggested:
The complete findings of this joint study were contained in a five-volume report issued by North American and submitted to MSC early in February 1964. (This document became known informally as the "Project Christmas Present Report.")
(29,000 feet) per second. Additional Details: here....
Those attending the meeting foresaw a number of problems:
Further analysis of the management system was necessary to determine changes needed in the checkout unit.
Flight profiles for Saturn IB missions for heatshield qualification purposes proved to be a little more difficult because "nobody would or could define the requirements or constraints, or test objectives." In other words, MSFC requirements for booster development test objectives and those of MSC for the spacecraft heatshield conflicted. So compromises had to be forged. Finally Ted H. Skopinski and other members of MPAD bundled up all of ASPO's correspondence on the subject generated from the various pertinent sources: MSFC, MSC, and contractors. From this, the Skopinski group drafted "broad term test objectives and constraints" for the first two Saturn IB flights (missions 201 and 202). Generally, these were to man-rate the launch vehicle and the CSM and to "conduct entry tests at superorbital entry velocities" (8,500 to 8,800 meters per second) (28,000 to 29,000 feet per second). Skopinski also enumerated specific test objectives covering the whole spacecraft-launch vehicle development test program. These were first distributed on March 27, and adjustments were made several times later in the year.
During this same week, Pratt and Whitney Aircraft tested a LEM-type fuel cell for 400 hours without shutdown and reported no leaks.
Three days later, Shea reported to the MSC Senior Staff that Apollo landings would be primarily on water. The only exceptions, he said, would be pad aborts and emergency landings. With this question of "wet" versus "dry" landing modes settled, Christopher C. Kraft, Jr., Assistant Director for Flight Operations, brought up the unpleasant problem of the CM's having two stable attitudes while afloat - and especially the apex-down one. This upside-down attitude, Kraft emphasized, submerged the vehicle's recovery antennas and posed a very real possibility of flooding in rough seas. Shea countered that these problems could be "put to bed" by using some type of inflatable device to upright the spacecraft.
The Block II design arose from the need to add docking and crew transfer capability to the CM. Reduction of the CM control weight (from 9,500 to 9,100 kilograms (21,000 to 20,000 pounds)) and deficiencies in several major subsystems added to the scope of the redesign. Additional Details: here....
Also MSC advised Grumman that estimates of the metabolic rates for astronauts on the lunar surface had been increased. The major effect of this change was an increase in the requirements for oxygen and water for the portable life support system.
For the first time, three representative Apollo space suits were used in the CM couches. Pressurized suit demonstrations, with three suited astronauts lying side by side in the couches, showed that the prototype suit shoulders and elbows overlapped and prevented effective operation of the CM displays and controls. Previous tests, using only one suited subject, had indicated that suit mobility was adequate. Gemini suits, tested under the same conditions, proved much more usable. Moreover, using Gemini suits for Apollo earth orbital missions promised a substantial financial saving. As a result of further tests conducted in May, the decision was made to use the Gemini suits for these missions. The existing Apollo space suit contract effort was redirected to concentrate on later Apollo flights. A redesign of the Apollo suit shoulders and elbows also was begun.
The astronauts' review was held on October 5 and 6. It included demonstrations of entering and getting out of the LEM, techniques for climbing and descending the ladder, and crew mobility inside the spacecraft. The general inspection was held on the 7th and the Review Board met on the 8th. Those attending the review used request for change (RFC) forms to propose spacecraft design alterations. Before submission to the Board, these requests were discussed by contractor personnel and NASA coordinators to assess their effect upon system design, interfaces, weight, and reliability.
The inspection categories were crew provisions; controls, displays, and lighting; the stabilization and control system and the guidance and navigation radar; electrical power; propulsion (ascent, descent, reaction control system, and pyrotechnics ; power generation cryogenic storage and fuel cell assemblies ; environmental control; communications and instrumentation; structures and landing gear; scientific equipment; and reliability and quality' control. A total of 148 RFCs were submitted. Most were aimed at enhancing the spacecraft's operational capability; considerable attention also was given to quality and reliability and to ground checkout of various systems. No major redesigns of the configuration were suggested.
As a result of this review, the Board recommended that Grumman take immediate action on those RFC's which it had approved. Further, the LEM contractor and MSC should promptly investigate those items which the Board had assigned for further study. On the basis of the revised M-5 configuration, Grumman could proceed with LEM development and qualification. This updated mockup would be the basis for tooling and fabrication of the initial hardware as well.
In another letter on October 16, the Project Office notified Grumman that no requirement existed for remote operation of either the rendezvous radar transponder or the stabilization and control system. The letter added, however, that the possibility of an incapacitated CSM astronaut must be considered and that for design purposes Grumman should assume that the astronaut would perform certain functions prior to becoming completely disabled. These functions could include turning on the transponder and the SCS. No CSM maneuvers would be required during the period in which the CSM astronaut was disabled but the CSM must remain stabilized during LEM ascent coast and rendezvous and docking phases.
It was acknowledged that a single mission could not serve to "completely define all the spacecraft functional requirements" but "such a mission has considerable value as a standard for various purposes on the Apollo Program."
Specifically, the DRM would be used for weight reporting, electrical power reporting, reliability modeling, engineering simulation, crew task analyses, mission-related Interface Control Documents, and trade-off studies.
Each group represented had a different interpretation of the reasons for the excessively high surface recession. The conclusion was that a second flight of the heatshield materials on the Scout would not particularly improve the understanding of the material's performance because of the limited variation in reentry trajectory and flight conditions obtainable with the Scout vehicle.
At about the same time, Grumman was analyzing the auxiliary battery requirements of the spacecraft. The contractor found that, under the worst possible conditions (i.e., lunar abort), the LEM would need about 1,700 watt-hours of auxiliary power. Accordingly, Grumman recommended one 1,700 watt-hour or two 850 watt-hour batteries (23 and 29.5 kg (50 and 65 lbs), respectively) in the spacecraft's ascent stage.
| Item | Mission Success | Crew Safety |
|---|---|---|
| Space suit | 0.9995 | 0.99991 |
| PLSS (Liquid cooled) | 0.9995 | 0.99999 |
| Complete assembly | 0.999 | 0.9999 |
Personnel from MSC, MSFC, KSC, OMSF, and North American attended the meeting. Included in the discussions were a review of the EDS design for both the launch vehicle and spacecraft along with related ground support equipment; a review of the differences of design and checkout concepts; and a review of EDS status lights in the spacecraft.
With reference to lighting, he said all numerics should be green, nomenclature and status lights white, and caution lights should be aviation yellow. All panel lighting should be dimmable throughout the entire range of brightness, including off.
In regard to nomenclature, Slayton pointed out that abbreviations on the DSKY should conform to the North American Interface Control Document (ICD). The referenced ICD was being reviewed by Grumman and North American and was scheduled to be signed December 1, 1964.
Referring to the caution and warning system, he pointed out that all caution lights on the DSKY should be gated into the primary navigation and guidance system (PNGS) caution light on the main instrument panel of both vehicles and into the PNGS caution light on the lower equipment bay panel of the CM.
Slayton requested that preliminary designs of the DSKY panel be submitted to the Subsystem Managers for Controls and Displays for review and approval.
At the same time, engineers at the Center began studying ways to increase the engine's thrust. Because of the LEM's weight gains, the engine must either be uprated or it would have to burn longer. Preliminary studies showed that, by using a phase "B" chamber (designed for a chamber pressure of 689.5 kilonewtons per sq m (100 psia)), thus producing chamber pressure of about 792.9 kilonewtons (115 psia), the thrust could be increased from 1,587 to 1,814 kg (3,500 to 4,000 lbs). Moreover, this could be accomplished with the present pressurization and propellant feed systems.