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At the third meeting of STG's New Projects Panel, Alan B. Kehlet presented suggestions for the multimanned reentry capsule. A lenticular-shaped vehicle was proposed, to ferry three occupants safely to earth from a lunar mission at a velocity of about 36,000 feet per second.
At a NASA staff conference at Monterey, Calif., officials discussed the advanced manned space flight program, the elements of which had been presented to Congress in January. The Goddard Space Flight Center was asked to define the basic assumptions to be used by all groups in the continuing study of the lunar mission. Some problems already raised were: the type of heatshield needed for reentry and tests required to qualify it, the kind of research and development firings, and conditions that would be encountered in cislunar flight. Members of STG would visit NASA Centers during April to define the tasks and request assistance. STG representatives were directed to maintain contact with the Centers and try to identify gaps in the technology. STG was also assigned the responsibility for preparing a first draft of specifications for a lunar spacecraft.
Robert O. Piland, Head of the STG Advanced Vehicle Team, and Stanley C. White of STG attended a meeting in Washington, D. C., sponsored by the NASA Office of Life Sciences Programs, to discuss radiation and its effect on manned space flight. Three consultants presented their views: John R. Winckler of the University of Minnesota, a cosmic-ray physicist; Cornelius A. Tobias of the University of California, a radiologist specializing in radiation effects on cells and other human subsystems; and Col. John E. Pickering, Director of Research at the Air Force School of Aviation Medicine. Their research showed that it would be impracticable to shield against the inner Van Allen belt radiation but possible to shield against the outer belt with a moderate amount of protection.
Three of the Apollo Technical Liaison Groups (Trajectory Analysis, Heating, and Human Factors) held their first meetings at the Ames Research Center.
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.
William W. Petynia of STG visited the Convair Astronautics Division of General Dynamics Corporation to monitor the Apollo feasibility study contract. A selection of the M-1 in preference to the lenticular configuration had been made by Convair. May 17 was set as the date for the final Convair presentation to NASA.
Command module heatshield requirements, including heating versus time curves, were established by NAA for several design trajectories. A computer program method of analyzing the charring ablation process had been developed. By this means, it was possible to calculate the mass loss, surface char layer temperature, amount of heat conducted through the uncharred ablation material and insulation into the cabin, and temperature profile through the ablator and insulation layers. In February, NAA determined that a new and more refined computer program would be needed.
Charles W. Frick, Manager of the MSC Apollo Spacecraft Project Office, together with Maxime A. Faget, Charles W. Mathews, Christopher C. Kraft, Jr., John B. Lee, Owen E. Maynard, and Alan B. Kehlet of MSC and George M. Low of the NASA Office of Manned Space Flight, visited NAA at Downey, Calif. This was the first monthly meeting of the Apollo design and review team to survey NAA's progress in various areas, including the Apollo spacecraft heatshield, fuel cells, and service module.
A purchase request was being prepared by NASA for wind tunnel support services from the Air Force's Arnold Engineering Development Center in the amount of approximately $222,000. These wind tunnel tests were to provide design parameter data on static stability, dynamic stability, pressure stability, and heat transfer for the Apollo program. The funds were to cover tests during June and July 1962. Approximately $632,000 would be required in Fiscal Year 1963 to fund the tests scheduled to December 1962.
At the monthly Apollo spacecraft design review meeting at NAA, MSC representatives recommended that NAA and Avco Corporation prepare a comprehensive test plan for verifying the overall integrity of the heatshield including flight tests deemed necessary, without regard for anticipated hunch vehicle availability.
NAA was directed by the Apollo Spacecraft Project Office at the monthly design review meeting to design an earth landing system for a passive touchdown mode to include the command module cant angle limited to about five degrees and favoring offset center of gravity, no roll orientation control, no deployable heatshield, and depressurization of the reaction control system propellant prior to impact. At the same meeting, NAA was requested to use a single "kicker" rocket and a passive thrust-vector-control system for the spacecraft launch escape system.
The control layout of the command module aft compartment was released by NAA. This revised drawing incorporated the new umbilical locations in the lower heatshield, relocated the pitch-and-yaw engines symmetrically, eliminated the ground support equipment tower umbilical, and showed the resulting repositioning of tanks and equipment.
After the determination of the basic design of the spacecraft sequencer schematic, the effect of the deployment of the forward heatshield before tower jettison was studied by NAA. The sequence of events of both the launch escape system and earth landing system would be affected, making necessary the selection of different sequences for normal flights and abort conditions. A schematic was prepared to provide for these sequencing alternatives.
Command module (CM) flotation studies were made by NAA, in which the heatshield was assumed to be upright with no flooding having occurred between the CM inner and outer walls. The spacecraft was found to have two stable attitudes: the desired upright position and an unacceptable on-the-side position 128 degrees from the vertical. Further studies were scheduled to determine how much lower the CM center of gravity would have to be to eliminate the unacceptable stable condition and to measure the overall flotation stability when the CM heatshield was extended.
Preliminary studies were made by NAA to determine radiation instrument location, feasibility of shadow-shielding, and methods of determining direction of incidence of radiation. Preliminary requirements were established for the number and location of detectors and for information display.
North American made a number of changes in the layout of the CM:
Representatives of North American, Langley Research Center, Ames Research Center, and MSC discussed CM reentry heating rates. They agreed on estimates of heating on the CM blunt face, which absorbed the brunt of reentry, but afterbody heating rates were not as clearly defined. North American was studying Project Mercury flight data and recent Apollo wind tunnel tests to arrive at revised estimates.
Two aerodynamic strakes were added to the CM to eliminate the danger of a hypersonic apex-forward trim point on reentry. [During a high-altitude launch escape system (LES) abort, the crew would undergo excessive g forces if the CM were to trim apex forward. During a low-altitude abort, there was the potential problem of the apex cover not clearing the CM. The strakes, located in the yaw plane, had a maximum span of one foot and resulted in significant weight penalties. The size of the strakes had to be increased later because of changes in the CM which moved the center of gravity forward and because of the additional ablative material needed to combat the increased heating of the strakes during reentry. Removal of the strakes would cause a major redesign to permit the apex cover to be jettisoned in the low angle-of-attack (apex forward) region. In the summer of 1963, however, MSC and North American representatives agreed that the strakes should be removed and an apex-mounted flap be added. The flap could be jettisoned with the LES tower during normal missions and retained with the CM during a LES abort.
North American then suggested a "tower flap dual mode" approach. This concept incorporated fixed surfaces at the upper end of the LES tower which would be exposed to the air stream after jettison of the expended rocket casing, For aborts below 9,140 meters (30,000 feet), the jettison motor would pull away the expended motor casing, the LES tower, and apex cover. The contractor carried out extensive wind tunnel tests of this configuration and reported to MSC during October that a 0.5941-square-meter (920-square-inch) planer flap located in the upper bay of the LES, coupled with a more favorable CM center of gravity, would be required to solve the reentry problem.
An independent investigation of deployable aerodynamic surfaces, or canards, at the forward end of the LES rocket motor was also being conducted. These canards would act as lifting surfaces to destabilize the LES and cause it to reorient the spacecraft to a heatshield-forward position.
NASA authorized North American to extend until June 10 the CM heatshield development program. This gave the company time to evaluate and recommend one of the three ablative materials still under consideration. The materials were subjected to tests of thermal performance, physical and mechanical properties, and structural compatibility with the existing heatshield substructure. North American sought also to determine the manufacturing feasibility of placing the materials in a Fiberglas honeycomb matrix bonded to a steel substructure.
North American completed a backup testing program (authorized by MSC on November 20, 1962) on a number of ablative materials for the CM heatshield. Only one of the materials (Avcoat 5026-39) performed satisfactorily at low temperatures. During a meeting on June 18 at MSC, company representatives discussed the status of the backup heatshield program. This was followed by an Avco Corporation presentation on the primary heatshield development. As a result, MSC directed North American to terminate its backup program. Shortly thereafter, MSC approved the use of an airgun to fill the honeycomb core of the heatshield with ablative material.
ASPO reported that a different type of stainless steel would be used for the CM heatshield. The previous type proved too brittle at cryogenic temperatures. Aside from their low temperature properties, the two metals were quite; similar and no fabrication problems were anticipated.
Ames Research Center performed simulated meteoroid impact tests on the Avco Corporation heatshield structure. Four targets of ablator bonded to a stainless steel backup structure were tested. The ablator, in a Fiberglas honeycomb matrix, was 4.369 millimeters (0.172 inch) thick in two targets and 17.424 millimeters (0.686 inch) thick in the other two. Each ablator was tested at 116.48 K (-250 degrees F) and at room temperature, with no apparent difference in damage.
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.
MSC and North American representatives discussed preliminary analysis of the probabilities of mission success if the spacecraft were hit by meteoroids. The contractor believed that pressurized tankage in the SM must be penetrated before a failure was assumed. To MSC, this view appeared overly optimistic. MSC held that, as the failure criterion, no debris should result from meteoroid impact of the SM outer structure. [This change in criteria would cost several hundred pounds in meteoroid protection weight in the SM and LEM.] North American thought that penetration of one half the depth of the heatshield on the conical surface of the CM was a failure. Here, MSC thought the contractor too conservative; full penetration could probably be allowed.
OMSF outlined launch vehicle development, spacecraft development, and crew performance demonstration missions, using the Saturn IB and Saturn V:
After completing estimates of the heating conditions for a series of MIT guided reentry trajectories, the MSC Engineering and Development Directorate recommended that the heatshield design philosophy be modified from the current "worst possible entry" to the "worst possible entry using either the primary or backup guidance mode." North American had drawn up the requirements early in 1962, with the intent of providing a heatshield that would not be a constraint on reentry. However, it was now deemed extremely unlikely that an entry, employing either the primary or backup guidance mode, would ever experience the heat loads that the contractor had designed for earlier. The ablator weight savings, using the MIT trajectories, could amount to several hundred pounds.
At Langley Research Center, representatives from Langley, MSC, Ames Research Center, Avco Corporation, and North American met to discuss their independent conclusions of the data gathered from the Scout test of the Apollo heatshield material and to determine whether a second test was advisable. Langley's report revealed that: the heatshield materials performed as predicted within the flight condition appropriate to Apollo; the excessive recession rates occurred during flight conditions which were more severe than those considered for the design of the heatshield or expected during Apollo reentries.
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.
During a mechanical loading test (simulating a 20-g reentry) the CM aft heatshield failed at 120 percent of maximum load. Structures and Mechanics Division engineers inspected the structure. They found that the inner skin had buckled, the damage extending three quarters of the way around the bolt circle that secured the heatshield to the spacecraft's inner structure. Their findings would be used along with data from the recent drop of boilerplate 28 to determine what redesign was necessary.
There appeared to be some confusion and/or disagreement concerning whether one or two successful Saturn V reentry tests were required to qualify the CM heatshield. A number of documents relating to instrumentation planning for the 501 and 502 flight indicated that two successful reentries would be required. The preliminary mission requirements document indicated that only a single successful reentry trajectory would be necessary. The decision would influence the measurement range capability of some heatshield transducers and the mission planning activity being conducted by the Apollo Trajectory Support Office. The Structures and Mechanics Division had been requested to provide Systems Engineering with its recommendation.
In its search for some method of reducing water impact pressures, North American was considering adding a 15- to 30.5-cm (6- to 12-in) "lump" to the CM's blunt face. The spacecraft manufacturer was also investigating such consequent factors as additional wind tunnel testing, the effect on heatshield design, and impact upon the overall Apollo program.
To solve the persisting problem of the integrity of the CM's aft heatshield during water impacts, MSC engineers were investigating several approaches: increasing the thickness of the face sheet (but with no change to the core itself); and replacing the stainless-steel honeycomb with a type of gridwork shell. Technicians felt that, of these two possibilities, the first seemed more efficient structurally.
Because of heat from the service propulsion engine (especially during insertion into lunar orbit), a serious thermal problem existed for equipment in the rear of the SM. Reviewing the rendezvous radar's installation, the Guidance and Control Division felt that a heatshield might be needed to protect the equipment. Similar problems might also be encountered with the steerable antenna.
The Structures and Mechanics Division (SMD) summarized the thermal status of antennas for the Apollo spacecraft (both CSM and LEM). Generally, most troubles stemmed from plume impingement by the reaction control or radiation from the service propulsion engines. These problems, SMD reported, were being solved by increasing the weight of an antenna either its structural weight or its insulation; by shielding it from the engines' exhaust; by isolating its more critical components; or by a combination of these methods.
Apollo boilerplate 28 underwent its second water impact test. Despite its strengthened aft structure, in this and a subsequent drop on February 9 the vehicle again suffered damage to the aft heatshield and bulkhead, though far less severe than that experienced in its initial test. The impact problem, it was obvious, was not yet solved.
North American gave boilerplate 28 its third water drop test. Upon impact, the spacecraft again suffered some structural damage to the heatshield and the core, though much less than it had experienced on its initial drop. Conditions in this test were at least as severe as in previous ones, yet the vehicle remained watertight.
As a result of the decision for an all-battery LEM, MSC advised Grumman that power for the entire pre- separation checkout of the spacecraft would be drawn from that module's batteries (instead of only during the 30 minutes prior to separation). This change simplified the electrical mating between the two spacecraft and obviated an additional battery charger in the CSM. From docking until the start of the checkout, however, the CSM would still furnish power to the LEM.
TWX, James L. Neal, MSC, to GAEC, Attn: R. S. Mullaney. April 30, 1965.
During the Month
Grumman reported two major problems with the LEM's descent engine:
Developmental testing began on a new landing device for the CM, one using rockets (mounted on the heatshield) that would be ignited immediately before impact. The current method for ensuring the integrity of the spacecraft during a landing in rough water involved strengthening of the aft structure. The new concept, should it prove practicable, would offer a twofold advantage: first, it would lighten the CM considerably; second, it would provide an improved emergency landing capability.
Crew Systems Division (CSD) representatives contracted with Northrop Space Laboratories to study physiological effects of tailward g forces. (CSD believed these forces might be "very hazardous." Consequently, the lowest impact limits for Apollo missions were in that direction.) Northrop would study bradycardia (slow heart rate) in animals induced by such acceleration, and would apply these findings to humans. CSD hoped thereby to determine whether current limits were "ultraconservative."
John H. Disher, Director of the OMSF Apollo Test Office, stressed two broad areas open to concern in the Apollo spacecraft heatshield development program:
MSC had planned to qualify the Block II heatshield by flight tests of modified Block I spacecraft 017 and 020. Some of the Block II changes could not be incorporated into modified Block I spacecraft in time to meet the current schedule and limitations of facilities would not permit full evaluation of all modifications by ground testing.
Disher suggested to Apollo Program Director Samuel C. Phillips that ASPO Manager Joseph Shea be asked to present physical descriptions of the Block I and Block II heatshields, and interim versions as applied to specific spacecraft, as well as the test plan that would ensure adequacy of heatshields to meet mission requirements. Memorandum, Disher to Phillips, "Apollo Spacecraft Heat Shield," June 28, 1965.
MSC directed North American to design the CM to store one integrated thermal meteoroid garment (TMG), rather than merely the thermal covering alone. The crewmen would carry the TMG into the LEM for use during extravehicular operations.
ASPO Manager Joseph F. Shea ordered Crew Systems Division to develop some type of protective devices that the astronauts might use to shield their eyes during a solar flare. ASPO regarded the risk of cataracts during these solar events as extraordinarily high. Although not mandatory, it was desirable that the crew could still see while wearing the devices. Should a flare occur while the crew manned the LEM, mission ground rules called for an abort back to the safety of the CSM; therefore, such devices would be needed for the CM alone.
North American conducted another in their series of impact tests with boilerplate 28. This drop tested the toroidal section of the spacecraft (heatshield and equipment bay structure) in impact at high angle and maximum horizontal velocity. The spacecraft suffered no visible damage. Some water leaked into the vehicle, but this was blamed on the boilerplate structure itself and the apex-down attitude after impact.
MSC's Assistant Director for Flight Operations, Christopher C. Kraft, Jr., told ASPO Manager Joseph F. Shea that postlanding operational procedures require that recovery force personnel have the capability of gaining access into the interior of the CM through the main crew hatch. This was necessary, he said, so recovery force swimmers could provide immediate aid to the crew, if required, and for normal postlanding operations by recovery engineers such as spacecraft shutdown, crew removal, data retrieval, etc.
Kraft said the crew compartment heatshield might char upon reentry in such a manner as to make it difficult to distinguish the outline of the main egress hatch. This potential problem and the necessity of applying a force outward to free the hatch might demand use of a "crow bar" tool to chip the ablator and apply a prying force on the hatch.
Since this would be a special tool, it would have to be distributed to recovery forces on a worldwide basis or be carried aboard the spacecraft. Kraft requested that the tool be mounted onboard the spacecraft in a manner to be readily accessible. He requested that the design incorporate a method to preclude loss of the tool - either by designing the tool to float or by attaching it to the spacecraft by a lanyard.
At a meeting with Grumman, MSC agreed with the contractor's basic design of the LEM's descent-stage base heatshield and its installation and access. MSC asked Grumman to demonstrate accessibility, installation, and removal of the heatshield on the M-4 mockup.
Preliminary results of the "fire-till-touchdown" study by Grumman indicated that this maneuver was not feasible. The engine might be exploded by driving the shock wave into the nozzles. The base heatshield temperature would exceed 1,789K (5,000 degrees F), which was high enough to melt portions of the structure, possibly causing destruction of the foot pads. The allowable pressure on the nonstructural elements of the base heatshield would be exceeded; and the descent engine flow field would tend to cause a "POGO" effect which would cause landing instability and could prevent engine cutoff.
As an outgrowth of the study, the landing probes would have to be made longer (137.1 to 187.9 cm [54 to 74 in] with automatic cutoff, 228.6 to 304.8 cm [90 to 120 in] with manual cutoff). The probe switches would be moved from the tip of the probe to the base, which was objectionable from the standpoint of a possible false reading due to probe dynamics.
Apollo-Saturn 201 was launched from Cape Kennedy, with liftoff of an Apollo Block I spacecraft (CSM 009) on a Saturn IB launch vehicle at 11:12:01 EST. Launched from Launch Complex 34, the unmanned suborbital mission was the first flight test of the Saturn IB and an Apollo spacecraft. Total launch weight was 22,000 kilograms.
Spacecraft communications blackout lasted 1 minute 22 seconds. Reentry was initiated with a space-fixed velocity of 29,000 kilometers per hour. CM structure and heatshields performed adequately. The CM was recovered by the USS Boxer from the Atlantic about 72 kilometers uprange from the planned landing point. (8.18 S x 11.15 W).
An Apollo Entry Performance Review Board was established by the MSC Director to review and validate the analytical tools as well as the Apollo operational corridor. The Board was set up because the performance of the ablation heatshield in the Apollo spacecraft, as then analyzed, imposed a limitation on the entry corridor at lunar return velocity. The following were named to the Board: Maxime A. Faget, MSC, chairman; Kenneth S. Kleinknecht, MSC; Eugene C. Draley and Don D. Davis, Jr., Langley Research Center; Alvin Seiff and Glen Goodwin, Ames Research Center; and Leo T. Chauvin, MSC, secretary.
Dale D. Myers, Apollo CSM Manager at North American Rockwell, wrote ASPO Manager George Low on the policy question of contractor and subcontractor support of the current Apollo flight program and potential follow-on activities. Support for such activities, Myers said, "can be seriously jeopardized if we permit . . . experienced, specialized personnel and unique facilities to become irretrievably lost to the program." He emphasized in particular the case of Aeronca, Inc., of Middletown, Ohio, manufacturer of stainless steel honeycomb panels that formed the structure of the CSM heatshield. Without some sort of sustaining activity, manufacturing skills and capabilities at Aeronca - and numerous other subcontractors and vendors - would rapidly wither. Myers earnestly solicited Low's views on the subject of subcontractor capability retention. In Low's response, he indicated that immediate action was being initiated to establish capability retention for the three most critical sources, Aeronca, Beech, and Pratt and Whitney, and a plan of action was being prepared for others.
Apollo Program Director Phillips asked ASPO Manager Low to hasten work on the study at North American to define reusability of systems aboard the CM. He asked Low for a review of the area in mid-February 1969 if sufficient data were available by then. Also, Phillips asked Low's recommendations for an effectivity date on any recovery operations to increase reusability of either spacecraft systems or of the complete vehicle. (North American submitted Space Division Report No. 69-463, dated August 29, 1969, recommending preflight preservation treatment and postflight refurbishment that could be accomplished on CMs and its components to enhance reusability. Removal of heatshield access ports and flushing with fresh water on the recovery ship was the only recommendation implemented, because the others were not judged cost effective.)