North American Aviation, Inc., selected the Aerospace Electrical Division of Westinghouse Electric Corporation to build the power conversion units for the command module (CM) electrical system. The units would convert direct current from the fuel cells to alternating current.

Four Navy officers were injured when an electrical spark ignited a fire in an altitude chamber, near the end of a 14-day experiment at the U.S. Navy Air Crew Equipment Laboratory, Philadelphia, Pa. The men were participating in a NASA experiment to determine the effect on humans of breathing pure oxygen for 14 days at simulated altitudes.

MSC ordered North American to provide batteries, wholly independent of the main electrical system in the CM, to fire all pyrotechnics aboard the spacecraft.

General Dynamics Convair completed structural assembly of the first launcher for the Little Joe II test program. During the next few weeks, electrical equipment installation, vehicle mating, and checkout were completed. The launcher was then disassembled and delivered to WSMR on April 25, 1963.

In support of NASA's manned space flight programs, Ames Research Center awarded a $150,000 contract to Westinghouse Electric Corporation for a one-year study of potential physiological damage in space caused by cosmic radiation.

North American announced that it had selected ITT's Industrial Products Division to provide battery chargers for the CSM, designed for an operational lifetime of 40,000 hours.

The first full-scale firing of the SM engine was conducted at the Arnold Engineering Development Center. At the start of the shutdown sequence, the engine thrust chamber valve remained open because of an electrical wiring error in the test facility. Consequently the engine ran at a reduced chamber pressure while the propellant in the fuel line was exhausted. During this shutdown transient, the engine's nozzle extension collapsed as a result of excessive pressure differential across the nozzle skin.

The first prototype of the CM battery for use during reentry was delivered to North American by Eagle-Picher Industries, Inc.

Because of the pure oxygen atmosphere specified for the spacecraft, North American reviewed its requirements for component testing. Recent evaluation of the CM circuit breakers had indicated a high probability that they would cause a fire. The company's reliability office recommended more flammability testing, not only on circuit breakers but on the control and display components as well. The reliability people recommended also that procurement specifications be amended to include such testing.

MSC authorized Grumman to proceed with procurement of a battery charger for the LEM, to replenish the portable life support system's power source. On the following day, Houston informed North American such a device was no longer needed in the CSM.

The Emergency Detection System (EDS) Design Sub-Panel of the Apollo-Saturn Electrical Systems Integration Panel held its first meeting at North American's Systems and Information Division facility at Downey, Calif. A. Dennett of MSC and W. G. Shields of MSFC co-chaired the meeting.

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.

ASPO approved the technique for LEM S-IVB separation during manned missions, a method recommended jointly by North American and Grumman. After the CSM docked with the LEM, the necessary electrical circuit between the two spacecraft would be closed manually. Explosive charges would then free the LEM from the adapter on the S-IVB.

Using boilerplate 14, North American simulated the mission for spacecraft 009. The test was conducted in two phases, with the vehicle on external and then internal power. All data showed satisfactory performance.

North American reported that qualification testing had been completed on two items of electrical hardware, the CSM battery charger and the pyrotechnic battery.

MSC notified Grumman that all electrically actuated explosive devices on the LEM would be fired by the Apollo standard initiator. This would be a common usage item with the CSM and would be the single wire configuration developed by NASA and provided as Government-furnished equipment.

Hamilton Standard Division was directed by Crew Systems Division to use a 2.27-kg (5-lbs) battery for all flight hardware if the power inputs indicated that it would meet the four-hr mission. The battery on order currently weighed 2.44 kg (5.4 lbs). This resulted in an inert weight saving of l.45 kg (3.2 lbs) and a total saving on the LEM and CSM of 5.44 kg (12 lbs).

The first test of the cryogenic gas storage system was successfully conducted from 12:30p.m. February 6 through 8:50 p.m. February 8 at the White Sands Test Facility (WSTF), N. Mex. Primary objectives were to demonstrate the compatibility between the ground support equipment and cryogenic subsystem with respect to mechanical, thermodynamic, and electrical interfaces during checkout, servicing, monitoring, and ground control. All objectives were attained.

The first integrated test of the service propulsion system, electrical power system, and cryogenic gas storage system was successfully conducted at the White Sands Test Facility.

The first manned flight of the Apollo CSM, the Apollo C category mission, was planned for the last quarter of 1966. Numerous problems with the Apollo Block I spacecraft resulted in a flight delay to February 1967. The crew of Virgil I. Grissom, Edward H. White II, and Roger B. Chaffee, was killed in a fire while testing their capsule on the pad on 27 January 1967, still weeks away from launch. The designation AS-204 was used by NASA for the flight at the time; the designation Apollo 1 was applied retroactively at the request of Grissom's widow.

Astronaut Frank Borman briefed the Apollo 204 Review Board after his inspection of the damaged command and service modules.

A main purpose of the inspection was to verify the position of circuit breakers and switches. In other major activities that day, the Pyrotechnic Installation Building was assigned to the Board to display the debris and spacecraft components after removal from Launch Complex 34; the Board began interviewing witnesses; and the Board Chairman asked NASA Associate Administrator for Manned Space Flight George E. Mueller for assistance in obtaining flame propagation experts to assist the Board. Experts might be obtained from Lewis Research Center, the Bureau of Mines, and the Federal Aviation Agency. The Board Chairman established an ad hoc committee to organize task panels to make the accident investigation systematically. The committee was composed of John J. Williams, KSC; E. Barton Geer, LaRC; Charles W. Mathews, NASA, Hq.; John F. Yardley, McDonnell Aircraft Corp.; George Jeffs, North American Aviation, Inc.; and Charles F. Strang, USAF.

The Service Module Disposition Panel (No. 21) report accepted by the Apollo 204 Review Board said test results had failed to show any SM anomalies due to SM systems and there was no indication that SM systems were responsible for initiating the January 27 fire.

Panel 21 had been charged with planning and executing SM activities in the Apollo 204 investigation, beginning at the time the Board approved the command module demate. The task was carried out chiefly by Apollo line organizational elements in accordance with a plan approved by the Board and identifying documentation and control requirements.

The panel's major activities had been:

- Demating the service module and service module-lunar module adapter from the launch vehicle and moving them to the Manned Spacecraft Operations Building.
- Inspecting the exterior and interior areas of the service module.
- Making detailed system tests of all service module systems that were mechanically or electrically connected to the command module at the time of the accident.

George Low requested William M. Bland, MSC, to take action on two recommendations made by MSC Director Robert R. Gilruth:

1. Take stereo color photos of all spacecraft areas before they were closed out. This procedure had been invaluable during the Apollo Review Board's activities at KSC, and the same technique, applied during the manufacturing process of current spacecraft, might help answer questions raised subsequent to the closeout of an area and thereby save time.
2. Make additional requirements for the use of cover plates over spacecraft wire bundles. Greater use of cover plates during manufacturing, test, and perhaps even flight would prevent damage during subsequent activities.

ASPO Manager George M. Low pointed out to MSC Director of Engineering and Development Maxime A. Faget that apparently no single person at MSC was responsible for spacecraft wiring. Low said he would like to discuss naming a subsystem manager to follow this general area, including not only the wiring schematics, circuitry, circuit-breaker protection, etc., but also the detailed design, engineering, fabrication, and installation of wiring harnesses.

Circuit breakers being used in both CSM and LM were flammable, MSC ASPO Manager George Low told Engineering and Development Director Maxime A. Faget. Low said that although Structures and Mechanics Division was developing a coating to be applied to the circuit breakers, such a solution was not the best for the long run. He requested that the Instrumentation and Electronics Systems Division find replacement circuit breakers for Apollo - ideally, circuit breakers that would not bum and that would fit within the same volume as the existing ones, permitting replacement in panels already built. On July 12 Low wrote Faget again: "In light of the work that has gone on since my May 5, 1967, memo, are you now prepared to propose the use of metal-jacketed circuit breakers for Apollo spacecraft? If the answer is affirmative, then we should get specific direction to our contractors immediately. Also, have you surveyed the industry to see whether a replacement circuit breaker is available or will be available in the future?" Low requested an early reply.

Grumman Aircraft Engineering Corp.'s method of building wiring harness for the lunar module was acceptable, George Low, MSC Apollo Spacecraft Program Office Manager, wrote Apollo Program Manager Samuel C. Phillips at NASA Hq. Low had noted on a visit to Grumman on May 9 that many of the harnesses were being built on two-dimensional boards. In view of recent discussions of the command module wiring, Low requested Grumman to reexamine their practice and to reaffirm their position on two-versus three-dimensional wiring harnesses.

In his May 31 letter to Phillips, Low enclosed Grumman's reply and said that, in his opinion, Grumman's practice was acceptable because

1. most wire bundles on the LM were much thinner than the CSM wiring bundles and were much more flexible;
2. portions of the LM harness were often fabricated on a three dimensional segment of the harness board; and
3. connectors were usually mounted on metal brackets with the proper direction and clocking.

ASPO Manager George M. Low issued instructions that the changes and actions to be carried out by MSC as a result of the AS-204 accident investigation were the responsibility of CSM Manager Kenneth S. Kleinknecht. The changes and actions were summarized in Apollo Program Directive No. 29, dated July 6, 1967.

A short circuit occurred during checkout of CSM 020 at North American, Downey, Calif.

External power batteries in parallel with the reentry batteries had indicated low power and were replaced. During preparations to continue the test, arcing was reported and emergency shutdown procedures were applied. Investigation was under way to determine the cause of the arcing. Initial indications were that at least 100 amps were imposed on a small portion of the spacecraft wiring, causing some damage to the spacecraft batteries.

Top NASA and North American Rockwell management personnel discussed flammability problems associated with coax cables installed in CMs. It was determined that approximately 23 meters of flammable coax cable was in CM 101 and, when ignited with a nichrome wire, the cable would burn in oxygen at both 4.3 and 11.4 newtons per square centimeter (6.2 and 16.5 pounds per square inch). Burning rates varied from 30 to 305 centimeters per minute, depending upon the oxygen pressure and the direction of the flame front propagation. The cable was behind master display panels, along the top of the right-hand side of the cabin, vertically in the rear right-hand corner of the cabin, in the cabin feed-through area, and in the lower equipment bay. The group reviewed the detailed location of the cable, viewed movies of flammability tests, examined movies of the results of testing with fire breaks, discussed possible alternatives, and inspected cable installations in CMs 101 and 104.

The following alternatives were considered:

1. Replace all coax cable.
2. Wrap all coax cable with aluminum tape.
3. Partially wrap the cable to provide fire breaks. Tests at North American indicated that a 102-millimeter segment of wrapped cable with four layers of aluminum foil would provide a fire break. MSC tests indicated such a fire break was not adequate for multiple cables.
4. Leave the installation as it was.

The following factors were considered in reaching a decision for spacecraft 101:

1. The wiring in that spacecraft had been completed for several months. All subsystems had been installed and protective covers had been installed. Complete replacement or complete wrapping of all coax cables would be time consuming; it might take as long as three months, when taking retest into consideration. Additionally, in spite of extreme care, complete replacement or wrapping might do considerable damage to the installed wiring, and even partial wrapping might cause damage in many areas.
2. The coax cable could not self-ignite under any conditions.
3. In most installations, the coax cable was a separate bundle and not part of other wire bundles. An exception was the feed-through area in the lower right-hand corner of the cabin, where the coax cable was intertwined with other wires. Although power cables existed in this area, these were not high-current-carrying cables.
4. A minimum number of possible ignition sources existed in the vicinity of the coax cables, and a complex series of events would be required to ignite the cable.

In view of these factors, decisions for spacecraft 101 were:

1. The cable would be flown essentially as installed. The only exception was that the vertical cable bundle in the right-hand corner of the spacecraft would be wrapped with layers of aluminum tape. Each cable in this bundle would be individually wrapped.
2. An analysis by North American would document all other wiring near the coax cable, including the wire size, functions, maximum currents carried, and degree of circuit-breaker protection.
3. All possible ignition sources near the coax cable would be documented.
4. Tests would be made in boilerplate (BP) 1250 to determine the effects of fire breaks inherent in the installation.

The installation in spacecraft 2TV-1 would not be changed. This decision was made fully recognizing that more flammable material remained in 2TV-1 than in 101. However, the burning rate of coax cable had been demonstrated as very slow, and it was reasoned that the crew would have sufficient time to make an emergency exit in the vacuum chamber from 2TV-1 long before any dangerous situations would be encountered.

Officials also agreed that coax cable in boilerplate 1224 would not be ignited until after the results of the BP 1250 tests had been reviewed.

CSM Manager Kenneth S. Kleinknecht asked the Manager of the Resident Apollo Spacecraft Program Office (RASPO) at Downey to inform North American Rockwell that MSC had found the suggestion that aluminum replace teflon for solder joint inserts and outer armor sleeves in Apollo spacecraft plumbing unacceptable because

1. the teflon insert was designed to give an interference fit to prevent the passage of solder balls into the plumbing;
2. an aluminum insert could not be designed with an interference fit for obvious reasons;
3. the aluminum insert was tested at the beginning of the program and found to be inferior to the teflon insert; and
4. the aluminum armor seal could not be used as a replacement for the outer armor sleeves because it did not eliminate the creep problem of solder.

ASPO Manager George M. Low advised top officials in Headquarters, MSFC, and KSC that he was recommending the use of 100 percent oxygen in the cabin of the LM at launch. MSC had reached this decision, Low said, after thorough evaluation of system capabilities, requirements, safety, and crew procedures. The selection of pure oxygen was based on several important factors: reduced demand on the CSM's oxygen supply by some 2.7 kilograms; simplified crew procedures; the capability for immediate return to earth during earth-orbital missions in which docking was performed; and safe physiological characteristics. All of these factors, the ASPO Chief stated, outweighed the flammability question. Because the LM was unmanned on the pad, there was little electrical power in the vehicle at launch and therefore few ignition sources. Further, the adapter was filled with inert nitrogen and the danger of a hazardous condition was therefore minimal. Also, temperature and pressure sensors inside the LM could be used for fire detection, and fire could be fought while the mobile service structure was in place. As a result, Low stated, use of oxygen in the LM on the pad posed no more of a hazard than did hypergolics and liquid hydrogen and oxygen.

Results of a joint MSFC-MSC review of functional interfaces between the launch vehicle and spacecraft for Apollo 7 were forwarded to NASA Hq.

(The review had originally been requested by the Apollo 7 Crew Safety Review Board, headed by John D. Hodge.) The two Centers had tackled the task by identifying all electrical wiring between payload and booster, the requirement for each wire, a verification that the circuits indeed satisfied requirements, and an evaluation of the adequacy of test and checkout procedures. Several months of investigation, reported Teir and Low, had uncovered no areas of concern. Definition and function of the CSM instrument unit were both accurate and valid and ensured flight readiness.

ASPO Manager George M. Low asked Rocco A. Petrone, Launch Operations Director at KSC, to set up a special task team to review all paperwork and to inspect visually all hardware, to ensure proper spacecraft deployment during the Apollo 8 flight.

Apollo 8 contained a novel set of mechanical and electrical interfaces (CSM, LTA-B lunar module dummy, launch adapter, and Saturn V vehicle), Low observed. Furthermore, concern about these complex interfaces had increased because one of the adapter panels on Apollo 7 had not opened properly. What Low - as well as MSC Director Robert R. Gilruth - desired foremost was to preclude repetition of another situation such as had occurred during the Gemini IX mission, when the shroud panels covering the Agena target vehicle had only partially deployed and had produced the "angry alligator" that forced cancellation of docking plans on that earlier flight.

George M. Low discussed the status of a fire detection system for Apollo in a memorandum to Martin L. Raines, reminding him that such a system had been under consideration since the accident in January 1967. Low said: "Yesterday, Dr. (Maxime A.) Faget, you, and I participated in a meeting to review the current status of a flight fire detection system. It became quite clear that our state of knowledge about the physics and chemistry of fire in zero gravity is insufficient to permit the design and development of a flightworthy fire detection system at this time. For this reason, we agreed that we would not be able to incorporate a fire detection system in any of the Apollo spacecraft. We also agreed that it would be most worthwhile to continue the development of a detection system for future spacecraft."

"Hey, we've got a problem here." The message from the Apollo 13 spacecraft to Houston ground controllers at 10:08 p.m. EDT on April 13, initiated an investigation to determine the cause of an oxygen tank failure that aborted the Apollo 13 mission.

The investigation terminated on June 15, when the Review Board accident report was released by NASA at a Headquarters press conference.

The Apollo 13 Review Board was established April 17 by George M. Low, NASA Deputy Administrator, and Thomas O. Paine, NASA Administrator, who appointed the Director of Langley Research Center, Edgar M. Cortright, as Review Board Chairman. On April 21 the members of the Board were named. In addition, by separate memos of April 20, the Aerospace Safety Advisory Panel was requested to review the procedures and findings of the Board and the Associate Administrator for Manned Space Flight was directed to provide records, data, and technical support as requested by the Board. The investigation indicated the accident was caused by a combination of mistakes and a somewhat deficient design. The following sequence of events led to the accident:

- After assembly and acceptance testing, the oxygen tank no. 2 that flew on Apollo 13 was shipped from Beech Aircraft Corp. to North American Rockwell (NR) in apparently satisfactory condition.

- However, the tank contained two inadequate protective thermostatic switches on the heater assembly, and they subsequently failed during ground test operations at Kennedy Space Center (KSC).

- In addition, the tank probably contained a loosely fitting fill tube assembly. This assembly was probably displaced during subsequent handling, which included an incident at the prime contractor's plant in which the tank was jarred.

- In itself, the displaced fill tube assembly was not particularly serious, but it led to improvised detanking procedures at KSC, which "almost certainly set the stage for the accident."

- Although Beech had not met any problem in detanking during acceptance tests, it was not possible to detank oxygen tank no. 2 using normal procedures at KSC. Tests and analyses indicate that the problem was gas leakage through the displaced fill tube assembly.

- The special detanking procedures at KSC subjected the tank to an extended period of heater operation and pressure cycling. "These procedures had not been used before, and the tank had not been qualified by test for the conditions experienced. However, the procedures did not violate the specifications which governed the operation of the heaters at KSC."

- In reviewing these procedures before the flight, officials of NASA, NR, and Beech did not recognize the possibility of damage from overheating. Many were not aware of the extended heater operation. In any event, adequate thermostatic switches might have been expected to protect the tank.

- A number of factors contributed to the presence of inadequate thermostatic switches in the heater assembly. The original 1962 specifications from NR to Beech Aircraft Corp. for the tank and heater assembly specified the use of 28-volt, direct-current power, which was used in the spacecraft. In 1965, NR issued a revised specification that stated the heaters should use a 65-volt dc power supply for tank pressurization; this was the power supply used at KSC to reduce pressurization time. Beech ordered switches for the Block II tanks but did not change the switch specifications to be compatible with 65-volt dc.

- The thermostatic switch discrepancy was not detected by NASA, NR, or Beech in their review of documentation, nor did tests identify the incompatibility of the switches with the ground support equipment (GSE) at KSC, "since neither qualification nor acceptance testing required switch cycling under load as should have been done. It was a serious oversight in which all parties shared."

- The thermostatic switches could accommodate the 65-volt dc during tank pressurization because they normally remained cool and closed. However, they could not open without damage with 65 volt dc power applied. They were not required to open until the special detanking. During this procedure, as the switches started to open when they reached their upper temperature limit, they were welded permanently closed by the resulting arc and were rendered inoperative as protective thermostats.

- Failure of the thermostatic switches to open could have been detected at KSC if switch operation had been checked by observing heater current readings on the oxygen tank heater control panel. Although not recognized at the time, the tank temperature readings indicated that the heaters had reached their temperature limit "and switch opening should have been expected."

- Subsequent tests showed that failure of the thermostatic switches probably permitted the temperature of the heater tube assembly to reach about 1,000 degrees F (810 K) in spots during the continuous eight-hour period of heater operation. Such heating had been shown by tests to damage severely the Teflon insulation on the fan motor wires near the heater assembly. "From that time on, including pad occupancy , the oxygen tank no. 2 was in a hazardous condition when filled with oxygen and electrically powered."

- Nearly 56 hours into the mission, the fan motor wiring, possibly moved by the fan stirring, short-circuited and ignited its insulation. Combustion in the oxygen tank "probably overheated and failed the wiring conduit where it entered the tank, and possibly a portion of the tank itself."

- The rapid expulsion of high-pressure oxygen which followed, "possibly augmented by combustion of insulation in the space surrounding the tank, blew off the outer panel to bay 4 of the SM, caused a leak in the high-pressure system of oxygen tank no. 1, damaged the high-gain antenna, caused other miscellaneous damage, and aborted the mission."

Based on the findings of the Board, a number of recommendations were made to preclude similar accidents in future space flights:

- The cryogenic oxygen storage system in the service module should be modified to:

- Remove from contact with the oxygen all wiring and unsealed motors that could potentially short-circuit and ignite adjacent materials; or otherwise ensure against an electrically induced fire in the tank.

- Minimize the use of Teflon, aluminum, and other relatively combustible materials in the presence of the oxygen and potential ignition sources.

- The modified cryogenic oxygen storage system should be subjected to a rigorous requalification program, including careful attention to potential operational problems.

- The warning systems on the Apollo spacecraft and in the Mission Control Center should be carefully reviewed and modified where appropriate, with specific attention to:

- Increasing the differential between master alarm trip levels and expected normal operating ranges to avoid unnecessary alarms.

- Changing the caution and warning system logic to prevent an out-of-limits alarm from blocking another alarm if a second quantity in the same subsystem went out of limits.

- Establishing a second level of limit sensing in Mission Control on critical quantities, with a visual or audible alarm that could not be easily overlooked.

- Providing independent talk-back indicators for each of the six fuel cell reactant valves plus a master alarm when any valve closed.

- Consumables and emergency equipment in the LM and the CM should be reviewed to determine whether steps should be taken to enhance their potential for use in a 'lifeboat' mode.

- MSC should complete the special tests and analyses under way to understand more completely the details of the Apollo 13 accident. In addition, the lunar module power system anomalies should receive careful attention. Other NASA Centers should continue support to MSC in the areas of analysis and test.

- Whenever significant anomalies occurred in critical subsystems during final preparation for launch, standard procedures should require a presentation of all prior anomalies on that particular piece of equipment, including those which have previously been corrected or explained. Critical decisions on flightworthiness should require the full participation of an expert "intimately familiar with the details of that subsystem."

- NASA should thoroughly reexamine all its spacecraft, launch vehicle, and ground systems containing high-density oxygen or other strong oxidizers, to identify and evaluate potential combustion hazards in the light of information developed in this investigation.

- NASA should conduct additional research on materials compatibility , ignition, and combustion in strong oxidizers at various gravity levels and on the characteristics of supercritical fluids. Where appropriate, new NASA design standards should be developed.

- MSC should reassess all Apollo spacecraft subsystems, and the engineering organizations responsible for them at MSC and at its prime contractors, to ensure adequate understanding and control of the engineering and manufacturing details at the subcontractor and vendor level. "Where necessary, organizational elements should be strengthened and in-depth reviews conducted on selected subsystems with emphasis on soundness of design, quality of manufacturing, adequacy of test, and operational experience."

To support the Apollo 13 Review Board, an MSC Apollo 13 Investigation Team, headed by Scott H. Simpkinson, was established with the several panels.

Spacecraft incident investigation, flight crew observations, flight operations and network ; photograph handling, processing, and cataloging ; corrective action study and implementation for the CSM, LM, and government-furnished equipment; related system evaluation; reaction processes in high-pressure fluid systems; high-pressure oxygen system survey; public affairs; and administration, communications, and procurement.

MSC Director Robert R. Gilruth reported MSC actions on the Apollo 13 Review Board recommendations.

- Fan motors had been removed from oxygen storage tanks in the service modules; the electrical leads had been encased in stainless steel sheaths with hermetically sealed headers and had been shielded from contact with the remaining Teflon parts.
- The modified cryogenic oxygen storage system had been subjected to a comprehensive recertification program developed in close coordination by North American Rockwell, Beech Aircraft Corp., and NASA. Requirements were founded on environmental as well as operational factors necessary to prove design capability.
- No major changes had been made in the caution and warning system.
- The LM and CSM consumables and emergency equipment had been reviewed to determine any design changes required to provide a safe return from lunar orbit in the event of a service module cryogenic-oxygen-supply loss. Three design changes were made in the CSM related to the oxygen tanks, an LM descent battery, and a water storage system in the CM.
- MSC had made special tests and analyses to understand the Apollo 13 accident better. The testing had reaffirmed the conclusions reached by the Apollo 13 Review Board.
- Significant anomalies in critical subsystems during final preparation for launch would be analyzed and resolved with authorized and documented corrective action in much the same manner as employed during the missions. An Apollo Program Directive for identification and resolution of significant failures and anomalies had been issued.
- A thorough reexamination of all spacecraft, launch vehicle, and ground systems containing high-density oxygen and other strong oxidizers was being made to identify and evaluate potential combustion hazards.
- Additional research was being conducted on materials compatibility, ignition, and combustion in strong oxidizers at various gravity levels and on the characteristics of supercritical fluids. Arc-ignition tests of the Apollo 14 oxygen-storage-system materials in both normal and overstressed modes indicated a positive margin of safety.
- MSC had organized a system-by-system task team effort and made comprehensive reassessments of each subsystem. Design and qualification of each subsystem was reaffirmed as adequate for current ground test and mission requirements with the exception of a heatshield blowout plug for dumping reaction-control-subsystem propellant for launch aborts.