Source: This document taken from the Report of Apollo 204 Review Board
NASA Historical Reference Collection, NASA History Division, NASA Headquarters, Washington, DC.

PART V. INVESTIGATION AND ANALYSES

1. INSPECTION AND DISASSEMBLY

Immediately after the accident additional security personnel were positioned at Launch Complex 34 and the Complex was impounded. Prior to disturbing any evidence numerous external and internal photographs were taken of the spacecraft. After crew removal, two experts entered the Command Module to verify switch positions. Small groups of NASA and North American Aviation management, Apollo 204 Review Board Members, Representatives and Consultants inspected the exterior of Spacecraft 012. On January 28, 1967 an astronaut entered the Command Module to verify additional switch positions needed to clarify data.

The Board established procedures for disassembly of Spacecraft 012. The first step of disassembly was to establish safe working conditions at the spacecraft. This was accomplished by:

  1. Removal of the Launch Escape System
  2. Removal or safetying of all pyrotechnics
  3. Examination of spacecraft structure for integrity
  4. Examination of all pressure vessels for potential hazards
  5. Sampling of spacecraft atmosphere for harmful contaminants

A series of close-up stereo photographs of the Command Module was taken to document the as-found condition of the spacecraft systems.

The task of searching the physical evidence was difficult and time-consuming because of the small entrance and confined area of the Command Module. In order to remove the components as quickly as possible, two persons at a time were permitted to enter the Command Module for component removal. After the removal of each component, photographs were taken of the exposed area. This step-by-step photography was used throughout the disassembly of the spacecraft (Enclosures 1 through 11). Approximately 5,000 photographs were taken.

After the couches were removed, a special false floor with removable 18-inch transparent squares was suspended from the existing couch strut fittings to provide access to the entire inside of the Command Module without disturbing evidence (Enclosure 12). A detailed inspection of the spacecraft interior was then performed followed by the preparation and approval by the Board of a Command Module disassembly plan. Command Module 014 was shipped to KSC on February 1, 1967 to assist the Apollo 204 Review Board in the investigation. This Command Module was placed in the Pyrotechnics Installation Building and was used to develop disassembly techniques for selected components prior to their removal from Command Module 012.

By February 7, 1967 the disassembly plan was fully operational. The concentrated effort of organized and coordinated component removal continued on a three-shift, seven-day-a-week basis. All suspect circumstances or conditions were brought to the attention of the Apollo 204 Review Board.

All interfaces such as electrical connectors, tubing joints, physical mounting of components, etc. were closely inspected and photographed immediately prior to, during and after disassembly. Each item removed from the Command Module was appropriately tagged, sealed in clean plastic containers and transported under the required security to bonded storage (Enclosure 13).

On February 17, 1967 the Board decided that removal and wiring tests had progressed to a point which allowed moving the Command Module without disturbing evidence. The Command Module was moved to the Pyrotechnics Installation Building at KSC where better working conditions were available.

With improved working conditions, it was found that a work schedule of two eight-hour shifts per day for six days a week was sufficient to keep pace with the analysis and disassembly planning. The only exception to this was a three-day-period of three eight-hour shifts per day used to remove the aft heat shield, move the Command Module to a more convenient work station and remove the crew compartment heat shield (Enclosures 14 and 15). The disassembly of the Command Module was completed on March 27, 1967.

From the beginning of disassembly, action was taken to catalog and place on display the hundreds of items that would be removed from the Command Module. The Pyrotechnics Installation Building was assigned to the Board for this purpose. A bonded storage room was established to receive and catalog components as they were removed. Command Module components were then displayed in a bonded area. The purpose of this area was to permit investigators to make visual examination of Command Module components (Enclosure 16). During the course of the disassembly over 1,000 items were removed from the Command Module. A list of all removed components was maintained and distributed weekly to the Board. This list identified the location of components in the Pyrotechnics Installation Building as well as those undergoing further analysis and tests at other locations.

Throughout the disassembly operation, experts meticulously studied the exposed portions of the Command Module. The relative consumption of combustibles and sooting patterns were studied for clues as to the site of the ignition source. All structural elements, covers and panels were examined for evidence of association with the ignition. Component systems and parts were studied inch by inch with magnifying glasses and frequently parts were taken into the laboratory for microscopic or metallurgical analysis. Wire bundles were given particular attention and after separation, the individual wires were examined under 7-power magnification for sites of possible arcing.

All components that showed evidence of abnormal fire effects were examined internally and many were tested for functionality. Many components showed burning of internal insulation or plotting material but in all cases they were exonerated on the basis of direction of flame travel or on the basis that there could be no communication with combustibles outside the component. Particularly suspect components were disassembled for detailed examination and analysis. All of the data developed by these visual and laboratory examinations were coordinated in making the final analysis as to probable ignition sources.

2. CHRONOLOGY OF THE FIRE

It is most likely that the fire began in the lower forward portion of the left-hand equipment bay. This would place the origin to the left of the Command Pilot, and considerably below the level of his couch.

Once initiated, the fire burned in three stages. The first stage with its associated rapid temperature rise and increase in the cabin pressure, terminated approximately 15 seconds after the verbal report of fire. At this time (about 23:31:19 GMT) the pressure vessel, which constitutes the Command Module cabin, ruptured. During this first stage of the fire, flames moved rapidly from the point of ignition, traveling along the Raschel net debris traps which were installed in the Command Module to prevent items from dropping into equipment areas during tests or flight. At the same time, Velcro strips positioned near the ignition point also burned.

Based upon pressure and temperature measurements taken during the fire, the fire was not intense until about 23:31:12 GMT. The slow rate of buildup of the fire during the early portion of the first stage is consistent with the view that ignition occurred in a zone containing little combustible material. The slow rise of pressure could also result from absorption of most of the heat by the aluminum structure of the Command Module. The original flames rose vertically and then spread out across the cabin ceiling. The debris traps provided not only combustible material and a path for the spread of the flames but also firebrands of burning molten nylon. The scattering of these firebrands contributed to the spread of the flames.

By 23:31:12 GMT the fire had broken from its point of origin. Evidence is strong that a wall of flames extended along the left wall of the module, preventing the Command Pilot, occupying the left hand couch from reaching the valve which would vent the Command Module to the outside atmosphere. Although operation of this valve, located on a shelf above the left hand equipment bay, is the first step in established emergency egress procedures, such action would have been to no avail because the venting capacity was insufficient to prevent the rapid build-up of pressure due to the fire. It is estimated that opening the valve would have delayed Command Module rupture by less than one second.

Emergency procedures called for the Senior Pilot, occupying the center couch, to unlatch and remove the hatch while retaining his harness buckled. A number of witnesses who observed the television picture of the Command Module hatch window during this stage of the fire discerned motion that suggests that the Senior Pilot was reaching for the inner hatch handle. The Senior Pilot's harness buckle was found unopened after the fire indicating that he initiated the standard hatch opening procedure. Data from the Guidance and Navigation System indicate considerable activity within the Command Module after the fire was discovered. This activity is consistent with movement of the crew prompted by proximity of the fire or with the undertaking of standard emergency egress procedures.

The Command Module is designed to withstand an internal pressure of approximately 13 pounds per square inch above external pressure without rupturing. Data recorded during the fire show that this design criteria was exceeded late in the first stage of the fire and that rupture occurred at about 23:31:19 GMT. The point of rupture was where the floor or aft bulkhead of the Command Module joins the wall, essentially opposite the point of origin of the fire. About three seconds before rupture, the final crew communication began at 23:31:16.8 GMT (a detailed discussion of the two voice transmissions during the fire is given in a subsequent section).

Rupture of the Command Module marked the beginning of the brief second stage of the fire. This stage is characterized by the period of greatest conflagration due to the forced convection that resulted from the outrush of gases through the rupture in the pressure vessel. The swirling flow scattered firebrands throughout the crew compartment spreading the fire. This stage of the fire ended at approximately 23:31:25 GMT. Evidence that the fire spread from the left hand side of the Command Module toward the rupture area was found on subsequent examination of the module. For example, the leg rest control handle on the left side of the left hand couch is fabricated from aluminum tubing. Tongues of flame pouring over the control handle melted its left side. However, a nylon button at the base of the handle was unconsumed and only slightly deformed. Similarly, flames spreading across the floor beneath the couches caused more burning on the left side of three nylon helmet covers than on the right. The underside of the couches above the helmet covers was relatively unsooted. A lack of soot indicates a fire of only short duration beneath the couches some time after the couch structure had become heated. Further, storage boxes situated on the floor were damaged only slightly. Fire across the floor of the spacecraft lasted but a few seconds and spread from left to right.

Damage to the crew suits is also indicative of the spread of the fire from left to right. The Command Pilot's suit was damaged worst while Senior Pilot's and Pilot's suits sustained progressively less damage.

Evidence of the intensity of the fire includes burst and burned aluminum tubes in the oxygen and coolant systems at floor level.

The pressure in the Command Module is estimated to have dropped to atmospheric pressure five or six seconds after rupture. The third and final stage of the fire began at about 23:31:25 GMT.

The third stage was characterized by rapid production of high concentrations of carbon monoxide. Following the loss of pressure in the Command Module and with fire now throughout the crew compartment, the remaining atmosphere quickly became deficient in oxygen so that it could not support continued combustion. Unlike the earlier stages where the flame was relatively smokeless, heavy smoke now formed and large amounts of soot were deposited on most spacecraft interior surfaces as they cooled. The third stage of the fire could not have lasted more than a few seconds because of the rapid depletion of oxygen. It is estimated that the Command Module atmosphere was lethal by 23:31:30 GMT, five seconds after the start of the third stage.

Although most of the fire inside the Command Module became extinguished shortly because of lack of oxygen, a localized, intense fire lingered in the area of the Environmental Control Unit. This unit is located in the left hand equipment bay, near the point where the fire is believed to have started. Failed oxygen and water/glycol lines in this area continued to supply oxygen and fuel to support the localized fire that melted the aft bulkhead and burned adjacent portions of the inner surface of the Command Module heat shield.

The loss of telemetry data at 23:31:22.4 GMT during the second phase of the fire makes determination of precise times of subsequent occurrences impossible. Thus all further times are based on less precise evidence such as entries in logs maintained by personnel monitoring various facets of the activity, review of voice tapes maintained of conversations between the Pad Leader and blockhouse monitors, and where so indicated, witness estimates.

3. DATA ANALYSIS

a. SCOPE OF INVESTIGATION

All data have been analyzed by the panels and the Board with frequent help from consultants and outside specialist groups. Specific tests of Spacecraft 012 equipment were initiated on approval by the Board where results would contribute to an understanding of the cause of the accident.

A summary of these results follows.

b. ANALYSIS OF RELEVANT TIME LINES

Enclosure 17 displays significant data that were obtained just prior to the report of the fire by the astronaut crew. These time lines cover the period of one minute before the fire report until all data signals were lost. The data shown includes signals from the gas chromatograph channel, the voltage of the AC Bus 2, the C-band beacon, the VHF telemetry carrier, the flow of oxygen into the suit loop, various indicators of spacecraft motion, the biomedical data from the Senior Pilot, and audio signals (voice and noise) received on the S-band communication link. An analysis of each item and a summary of their correlation follows.

(1) Gas Chromatograph Telemetry Data Anomaly

The gas chromatograph was not installed for the Plugs-Out Test and the connector that carried the telemetry data signals and the required AC power was open ended and was placed on the gas chromatograph shelf prior to the test. Power to the AC in the connector was turned on during the test as required by the test plan.

A careful examination of the data records disclosed activity on this channel eight times up to and including the activity shown at approximately 23:30:50 GMT. Subsequent testing has demonstrated that the telemetry data lead in the connector has the characteristics of an antenna, and consequently can detect changes in electromagnetic fields within the spacecraft. Movement of this cable within a constant electromagnetic field will also produce signals of the magnitude observed during the Spacecraft 012 accident.

The disturbance at approximately 23:30:50 GMT indicates that such a change in the electromagnetic field took place. This change could have resulted from movement of the connector. Evidence indicates that although the connector was not in its originally stored position after the accident, it probably was there during the initial stages of the fire.

(2) AC Bus 2 Voltage Anomaly

A momentary increase in AC Bus 2 voltage on all three phases was noted at approximately 9 seconds before the report of fire, and at the same time telemetry data from equipment powered from AC Bus 2 showed abnormalities. These were:

  1. Dropout of C-band decoder and transmitter outputs for 1.7 seconds.
  2. Momentary dropout of VHF-FM transmitter.
  3. Fluctuation of rotation controller null outputs.
  4. Gas chromatograph telemetry signal transient.

Other equipment connected to AC Bus 2 at this time had no data monitoring capability that would detect effects of power transients.

The power distribution system was in the standard configuration at the time of the anomaly. DC bus A was receiving power from the ground DC "A" power supply. This power supply in turn powered AC Bus 1 through inventor no. 1. Similarly DC Bus B received power from the DC "B" power supply and powered AC bus 2 through inverter no. 2.

A possible explanation for dropout of the C-band decoder and transmitter, the interruption of the VHF-FM transmitter and rise in AC Bus 2 voltage follows. The post-landing bus supplies power through a single conductor and circuit breaker to the power relay holding coils for both the C-band beacon and the VHF-FM transmitter. Temporary loss of voltage to the relay holding coils by unknown cause, would temporarily interrupt power to the C-band decoder and VHF-FM transmitter. The resulting transient to the voltage level on AC Bus 2 could account for other measured phenomena.

The most probable cause of the AC Bus 2 transient and associated indications was a momentary short or interruption of DC Bus B. Analysis and subsequent testing correlate with this conclusion as follows:

(a) AC Bus Transient

This high voltage indication can be interpreted as evidence of a momentary drop of DC voltage input to the inverter which results in a drop in AC output and a subsequent overshoot upon recovery. First indication of a disturbance was noted during apparent recovery. The voltage decrease was not seen because the channel was sampled only 10 times a second.

(b) C-band Beacon Dropout

The 1.7 second dropout observed is the minimum recovery time of the protective circuit internal to the beacon. A momentary interruption of AC Bus 2 power for a period as short as 10 milliseconds would cause the C-band beacon dropout. These results were verified by special tests on a C-band beacon similar to the one used in Spacecraft 012. The most probable cause of the beacon dropout was a momentary loss of AC input power to the beacon particularly since the transponder dropout was coincident with a transient on the AC Bus 2 and the beacon performed normally after recovery from the dropout unit loss of data.

(c) VHF-FM Transmitter Signal Dropout

The RF carrier dropout was observed by all monitoring ground stations and the duration of the dropout was approximately 20 milliseconds. The recorded data wave train from the VHF-FM transmitter also indicated dropout. A dropout of this nature has been duplicated by several special tests with a similar transmitter under similar conditions. Because the VHF transmitter recovered, the most probable cause of the dropout was a momentary interruption of the AC input power.

(d) Rotation Controller Null Output Transients

Momentary transients were noted on each of the three control axes. The rotation controller, whose output was reading slightly off null just prior to the anomaly (the controller was pinned), was supplied by phase A of AC Bus 2. Transient voltages on the phase A Bus would most likely be detected on the controller output. Special tests have shown that the null output transients experienced can be duplicated by a momentary interruption of AC Bus 2 phase power.

(e) Gas Chromatograph Telemetry Signal Transient

As previously discussed this transient could result from a change in the electromagnetic field. Such a change in the electromagnetic field could also be the result of electric arcing.

(3) Anomalies in Oxygen Flow

Enclosure 18 is a schematic of the suit loop. Oxygen is normally supplied from the surge tank and the service module cryogenic storage tanks through an oxygen regulator which controls the supply pressure to approximately 100 psi. An oxygen flow transducer is installed in the supply line downstream of the oxygen regulator and oxygen is supplied to the suit loop through a demand regulator. The oxygen in the suit loop is circulated to the three astronauts through three separate branches. Each branch has an individual flow rate transducer.

The flow rate of oxygen started to increase approximately 40 seconds before the reported fire. The output limit or saturation of the flow transducer, which corresponds to a flow of 1.033 pounds per hour, was reached approximately 5 seconds before the first fire report. The oxygen flow transducer stayed in this saturated condition until loss of data occurred. The Caution and Warning Alarm was actuated 15 seconds after the oxygen flow exceeded one pound per hour. This delay is normal and prevents actuation of the Caution and Warning Alarm during normal short duration, high flow conditions.

The stable oxygen surge tank pressure, coupled with the normal oxygen regulated pressure, indicates that oxygen flow rate was not greater than 3 to 6 pounds per hour until approximately 7 seconds after the first fire report. Beyond that time the oxygen flow rate was much higher.

The initial flow rate increase is probably due to crew movement which normally results in increased leakage to the cabin at low differential pressure conditions. Enclosure 19 shows that at approximately 23:31:03 GMT the oxygen flow had increased to the suit loop to the extent that the pressure differentials across the suits and compressor were increasing. There was an indication that suit circuit flow through the Senior Pilot's suit was interrupted for about two seconds at approximately 23:31:09 GMT. This interruption in flow to the Senior Pilot's suit is not completely understood; however, it probably was caused by manipulation of the suit hoses or associated controls.

(4) Indications of Spacecraft Motion

A number of individual signals were received which are indicative of slight motions of the spacecraft within the last minute prior to the first fire report. These signals were of a random nature and are similar to signals that were obtained from the spacecraft during known crew movement.

These signals included corrective torque signals to the gyrocompasses in the Inertial Measuring Unit, a brushing or tapping of the Command Pilot's live microphone, the previously mentioned increase in oxygen flow attributed to suit leakage and an increase in the attention level of the Senior Pilot most noticeable between 23:30:30 GMT and 23:30:45 GMT. The Senior Pilot was the only member of the crew for whom biomedical data were recorded.

The nature of activity of the crew during this period could not be determined.

(5) S-Band Transmissions

There were no voice transmissions from the spacecraft from 23:30:14 GMT until the first indication of fire in the spacecraft by the crew. During this time the Command Pilot had a live microphone condition as noted previously. Two voice transmissions were subsequently received. The first of these was the first indication of the existence of a fire by the crew.

  • The Live Microphone Anomaly

    Voice tape analysis and instrumentation data records show that a live microphone, constant-keying condition, existed from the Command Pilot position during a considerable portion of the final test period. This condition apparently did not exist beyond the first of the final two voice transmissions from the spacecraft.

    Audio circuits are normally actuated by a crewman pressing his Push-to-Talk (PTT) button on the cobra cable or in the Command Pilot's case by pressing his controller PTT or his cobra cable PTT button (Enclosure 20). This action serves to ground the microphone amplifier in the individual crewman's audio panel as well as the diode gate in the audio center on the S-band audio output. These functions allow signals to modulate the S-band transmitter.

    The problem has been isolated to the PTT or keying line that runs between the cobra cable, translation controller, Command Pilot audio control panel and the audio center. Crew attempts to isolate the problem were unsuccessful although the Command Pilot's cobra cable was absolved after troubleshooting. Subsequent testing has also failed to disclose the cause of this problem.

    Power limitations and subsequent testing of this circuitry indicates that sufficient current cannot be carried by this keying circuitry for it to be considered a possible ignition source.

  • Voice Transmission

    The final two voice transmissions were made on S-band. No voice communications on VHF were made from the spacecraft during this period. The first transmissions lasted from 23:31:04.7 GMT through 23:31:10 GMT and the second lasted from 23:31:16.8 GMT through 23:31:21.8 GMT. The tape recordings of these transmissions have been analyzed extensively and the results are presented subsequently.

    • (6) Cabin Pressure Rise

    The cabin pressure for the period from first report of the fire through loss of signal is shown in Enclosure 21.

    First indication by either the cabin pressure or battery compartment (open to the cabin) sensors of a pressure increase occurred at approximately 23:21:11 GMT or about 6 seconds after the crew first reported the fire. The pressure exceeded the range of these transducers, 17 pounds per square inch absolute (psia) for the cabin and 21 psia for the battery compartment transducers by 23:31:16 GMT. Data from this time until loss of signal were derived from the response of Guidance and Navigation equipment to the different pressure changes. The cabin ruptured at a time of about 23:31:19 GMT and at a pressure of at least 29 psia.

    Rupture occurred in the -Y, +Z quadrant and the resulting jet of hot gases caused extensive damage to the exterior structure (Enclosures 22 through 25).

    (7) Summary of Relevant Events

    Between 30 and 45 seconds prior to the report of fire, both the Command Pilot and Senior Pilot were active. The nature and level of the activity remain unknown. Except for the transients in data measurements that occurred approximately 9 seconds prior to the report of the fire, there are no other identified relevant events that preceded the fire. It should be noted that these data transients and subsequent activity of the crew may as easily be associated with the result of the fire as with the cause.

    The increase in oxygen flow to the suit loop prior to and immediately following the report of the fire and its effect on the pressure distribution within the suit loop is the result of normal demand regulator response to oxygen leaking from the circuit to the cabin. This is further compounded by the response of the regulator to the rise in cabin pressure.

    (c) ANALYSIS OF CREW VOICE TRANSMISSION DURING THE FIRE

    The tape transients of the voice tapes from the Command Module during the period of the fire have been analyzed extensively. These analyses included a review of all transmissions prior to the fire that were made by the crew during the test in an attempt to aid in the determination of who made these last two transmissions and what was said. These analyses were made by NASA personnel familiar with the communication systems, the crew and their voice characteristics, the sequence of events before, during and after the fire as determined during the investigation. The Board also reviewed these transmissions. Experts at the Bell Telephone Laboratories performed extensive analyses of the tape record.

    Except for a portion of the first transmission which is quite clear, the remainder of the transmissions are not clear and it is impossible to define exactly what was said by the crew.

    Two points made by the Bell Telephone Laboratory experts should be noted:

    1. The present state-of-the-art of analysis of voice records is such that little if anything can be determined as to what was said if the recording is not sufficiently clear to be intelligible by listening alone. Analysis can provide some clues as to who may have made the transmissions; however, these clues are not definitive.

    2. When the recording of the transmission is not clear, there will be nearly as many interpretations of what was said as there are qualified listeners.

      A summary of various interpretations of these transmissions is made in the following paragraphs.

      The analysis of the first transmission is as follows:

      This transmission began at 23:31:04.7 GMT with an exclamatory remark. This transmission is not clear. Listeners believe this initial remark was "Hey " or "Fire" but this is not certain.

      Some listeners believe and laboratory analysis supports this belief that this transmission was made by the Command Pilot. This remark is followed by a short period of noise (bumping sound, etc.).

    The second portion of this first transmission begins at 23:31:06.2 GMT with unclear word. Most listeners believe the first word to be one of the following:
    • "I've"
    • "We've"

    The remainder of this transmission is quite clear and is: "...Got a fire in the cockpit," followed by a clipped word sounding like "Uheh," which ended at 23:31:10 GMT. Many listeners believed this transmission was made by the Pilot and laboratory analysis tend to support this belief. However, no firm conclusion can be drawn.

    The analysis of the second transmission is as follows:

    Following a 6.8-second period of no transmission, the second transmission began at 23:31:16.8 GMT and ended at 23:31:21.8 GMT. The entire second transmission is garbled and is, therefore, subject to wide variation of interpretation as to content and as to who made the transmission and no definitive transcription is possible.

    The general content of this transmission consists of what appears to be three separate phrases. It has been interpreted several ways by many listeners. The following is a list of some of the interpretations that have been made:

    1. "They're fighting a bad fire - Let's get out ....Open 'er up."
    2. "We've got a bad fire - Let's get out ....We're burning up."
    3. "I'm reporting a bad fire ....I'm getting out ...."

    This transmission ended with a cry of pain. Some listeners believe this transmission was made by the Pilot.

    It should be noted that:

    1. The total duration of these two transmissions was brief, lasting 10.3 seconds; the first lasted 5.3 seconds and the second lasted 5.0 seconds, with a 6.8-second period of no transmission.
    2. The transmissions provide evidence only of the time the crew first reported the existence of the fire and do not provide any information as to the cause of the fire.

    d. MEDICAL ANALYSIS

    Loss of consciousness was due to cerebral hypoxia due to cardiac arrest resulting from myocardial hypoxia. Factors of temperature, pressure and environmental concentrations of carbon monoxide, carbon dioxide, oxygen and pulmonary irritants were changing extremely rapidly. It is impossible to integrate these variables on the basis of available information with the dynamic physiological and metabolic conditions they produced, in order to arrive at a precise statement of time when consciousness was lost and when death supervened. The combined effect of these environmental factors dramatically increased the lethal effect of any factor by itself. It is estimated that consciousness was lost between 15 and 30 seconds after the first suit failed. Chances of resuscitation decreased rapidly thereafter and were irrevocably lost within 4 minutes.

    4. CAUSE OF THE APOLLO 204 FIRE

    The fire in Apollo 204 was most probably brought about by some minor malfunction or failure of equipment or wire insulation. This failure, which most likely will never be positively identified, initiated a sequence of events that culminated in the conflagration.

    A great deal of effort has been expended in an attempt to find this specific initiator. Although unsuccessful in this search, this effort has produced a fairly good understanding of the types of things that may have been the initiator and the types of things that probably could not have been the initiator.

    Electrostatic discharge, spontaneous combustion of flammable material, mechanically produced heat by machinery and heat from the impact of a struck object have been eliminated as reasonable possibilities of ignition of the fire. The flow of oxygen through orifices or metering valves can create heat through the excitation of resonating frequencies in the gas. However, a thorough examination of the hardware and evaluation of recorded performance of the equipment eliminates the energy of flowing oxygen as a possible initiator.

    The most obvious source of energy needed to initiate the fire existed in the spacecraft's power distribution system. Current carrying wires were distributed throughout every major region of the Command Module. The most likely ways in which electrical power can initiate a fire are the following:

    1. Through malfunction of the equipment being powered which in turn ignites or initiates a fire in nearby combustibles.
    2. Overload in the conductor resulting from shorts in equipment or wiring. This overload will cause the conductor to overheat and ignite nearby combustibles (Enclosure 26).
    3. Electric arcs that are created when the insulation is defeated between power carrying conductors and the spacecraft structure or equipment.

    A large majority of the wires were left undamaged. However, there were a number of cases where exposed wire showed extensive burning, overheating or complete destruction. There were also several places where pitting of exposed conductors and adjacent structure indicate that an electric arc had occurred.

    • MALFUNCTION OF ELECTRICAL POWERED EQUIPMENT

      After removal from the spacecraft, each component or subassembly was critically examined to determine whether or not it could be associated with the initiation of the fire. The vast majority of these could be classified as non-initiators on the basis of external examination and recorded performance. If, however, there was any suspicion that an item was involved with the initiation of the fire, it was subjected to intensive scrutiny that involved one or more of the following procedures:

      1. Laboratory analysis of damage, electrical continuity and resistance tests.
      2. Functional performance using established procedures for "bench checks."
      3. Careful disassembly which included repeating some of the above steps on individual parts of the assembly.

      The results of this effort led to the conclusion that none of the electrically powered spacecraft systems or subassemblies was associated with the initiation of the fire.

    • ELECTRICALLY OVERLOADED CONDUCTORS

      The Apollo spacecraft wiring is protected with Teflon insulation. Teflon was chosen as the insulating material after a series of tests clearly showed that it was the least likely to burn when overheated by shorting. Individual conductors in a wire bundle using Teflon-insulated wires could be melted to destruction without initiating a sustained fire in the bundle when located in a 100-percent oxygen atmosphere at 5 psia. The Teflon-insulating material provided a high degree of fire protection to wire bundles which may contain electrically overloaded wires. Primary protection to wiring in the spacecraft, however, was provided by circuit breakers and fuses which protected all power-carrying conductors. Critical analysis of all circuit breaker installations showed that this protection was provided adequately with only a few exceptions. Several indications of shorted wiring were made the subject of individual detailed investigations. These investigations have all proved negative except for a few cases that could not be exonerated completely.

    • ELECTRIC ARCS

      Teflon has excellent fire resistance but low resistance to cold flow. The Teflon covering on the wire used in Apollo 204 could be damaged easily or penetrated by abrasion. The covering could also be damaged when forced against the structure by poor installation. The Board found numerous examples in the wiring of poor installation, design and workmanship (an example is shown in Enclosure 27 where a wrench socket was found in the spacecraft.). If a power conducting wire experiences penetration of its insulation by the metal structure of the spacecraft or spacecraft components, an instantaneous short to ground is created at the point of conductor contact. An arc or a series of arcs between conductor and structure results. The arcing action may be terminated by the blowing away of molten metal at the point of contact, or if sufficient mechanical pressure exists, fusion between the conductor and structure may occur to create a continuous short. The previous occurrence of an arc can be determined through examination of hardware because a characteristic pit or crater is left at the location of contact. Tests in a 16.5 psia oxygen atmosphere have shown that sparks blown from arcs can ignite combustible material several inches from the arc. Circuit breakers and other practical circuit interrupting devices cannot act rapidly enough to prevent an arc. Thus, arcs cannot be eliminated as a potential source of ignition energy. As noted previously, there were strong data indications of an abrupt, short-duration voltage decrease. This is consistent with a quickly terminated arc. During the examination of hardware and wiring, particular emphasis was placed on locating craters near power cables. While several such craters were found, only one appeared to be linked closely to the time of the fire by other supporting evidence. A complete investigation of the evidence associated with this possible ignition source has relegated it to a low probability. Studies of fire damage patterns indicate that the most likely region for the start of the fire is underneath the lithium hydroxide access door. Damage is so extensive in this location that the physical evidence remaining provides little interpretive information (Enclosure 11). Power cable insulation passing under this door was potentially vulnerable to abrasion from the corner along the lower edge of the door. (Enclosure 28 shows this cable as it is installed on Spacecraft 014.) If this cable were the cause, it cannot be proven since both the power cable and the inside edge of the door were completely destroyed. It is most probable that the fire was initiated by an electric arc either in this location or in some other region near the Environmental Control Unit. Other powered cable in the Environmental Control Unit may have been the source but extensive destruction of them precludes a positive determination. The time of initiation probably coincides with the spacecraft power interruption at 23:30:55 GMT.

      The Board's investigation was facilitated by the wide application of simulation techniques. The consequences of several types of electrical faults were studied in this way. The most valuable simulation, however, employed full-scale fire tests in a boilerplate mock-up with combustibles arranged in the configuration of Command Module 012. These tests were conducted by KSC for the Board. Ignition was obtained by a hot wire in the general area in which ignition is suspected to have occurred in the fire and provisions were made for simulating rupture at proper time and location.

      Total time of the fire and the pressure history reproduced those of the fire quite closely thus adding confidence to the deduced origin and mode of propagation of the fire. Such simulation techniques should be applied in examining the fire hazards of future spacecraft. They provide a reliable means of assessing fire hazards. They have also demonstrated that laboratory tests on small samples may give misleading results.

    • EFFECT OF COOLANT ON ELECTRICAL WIRES AND EQUIPMENT

      The discussion of possible electric power distribution malfunctions in Apollo 204 cannot be complete without inclusion of the effect of Environmental Coolant System coolant leakage. The Apollo Block I Spacecraft uses RS-89 as a coolant. This coolant is a mixture of 62.5 percent ethylene glycol, 35.7 percent water, and 1.8 percent stabilizer and corrosion inhibitor. Although the mixture is not highly combustible, leakage and spillage of this fluid present a considerable fire hazards. The water evaporates more readily than the ethylene glycol and the inhibitor consists of two combustible salts which do not evaporate. Consequently, spilled coolant can become a dangerous combustible if it is not removed properly. The inhibitor mixture presents a second hazard in that it is also hygroscopic and electrically conductive. Thus, the residue from coolant that was spilled and subsequently evaporated, remained slightly wet. This residue is corrosive and may conduct electricity if it wets electrical wiring or equipment that does not have water-proof insulation. The conductive path so formed will progressively improve itself as dendrites grow through electrolytic action. The RS-89 coolant is particularly dangerous in the presence of damaged or improperly insulated electrical equipment and harnesses. During the design of Apollo Block I Spacecraft a decision was made to seal electrical components and connectors. As a result, many of the Spacecraft 012 electrical systems were watertight (Block II Spacecraft are designed to have complete sealing of electrical equipment and harnesses).

      Coolant in the spacecraft is used to extract heat from the cabin atmosphere and from the circulation loop to the spacesuits. It also provides direct cooling through coldplates to numerous pieces of electrically powered equipment. Thus, the cooling system is extensive throughout the Command Module. The plumbing that carries the coolant is assembled from aluminum tubing utilizing both metallurgical (soldered, brazed or welded) and mechanical joints. Numerous plumbing designed in that strength margins were inadequate to resist damage from unplanned loads. Such loads may result during equipment installation or when tubing is used as hand-holds or is bumped by technicians working in the Command Module. The result was that a number of leaks in solder joints were experienced during the history of all Block I spacecraft. The mechanical joints also had leakage problems.

      There is no substantial evidence that coolant was involved in the initiation of the fire. However, this coolant, when spilled on damaged electrical wires and equipment, provides both the fuel and the ignition mechanism to start a fire. This has been demonstrated in laboratory tests.

    • SPACECRAFT ATMOSPHERE

      The use of pure oxygen in American spacecraft has been the subject of much consideration. The use of a diluent gas, either nitrogen or helium, in large proportions would undoubtedly reduce the risk of fire to a significant degree. At the same time it would introduce other operational problems and risks. There is no obvious advantage of one diluent over the other, although much progress has been made in developing the complex technology required for controlling gas concentrations to maintain a proper mixture reliably. This technology is still far from being fully developed. Furthermore, there are many difficult operational problems that must be solved in a reliable manner in order to decrease rather than increase the risks before undertaking the use of a two-gas system.

    • SUMMARY

      Although the Board was not able to determine conclusively the specific initiator of the Apollo 204 fire, it has identified the conditions which led to the disaster. These conditions were:

      1. A sealed cabin, pressurized with an oxygen atmosphere.
      2. An extensive distribution of combustible materials in the cabin.
      3. Vulnerable wiring carrying spacecraft power.
      4. Vulnerable plumbing carrying a combustible and corrosive coolant.
      5. Inadequate provisions for the crew to escape.
      6. Inadequate provisions for rescue or medical assistance.

      Having identified the condition that led to the disaster, the Board addressed itself to the question of how these conditions came to exist. Careful consideration of this question leads the Board to the conclusion that in its devotion to the many difficult problems of space travel, the Apollo team failed to give adequate attention to certain mundane but equally vital questions of crew safety. The Board's investigation revealed many deficiencies in design and engineering, manufacture and quality control. When these deficiencies are corrected the overall reliability of the Apollo Program will be increased greatly.

    History Of The Accident | Chronology From T-10 Minutes | Description of Test Sequence And Objectives | Findings, Determinations And Recommendations

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    Updated August 4, 2006
    Steve Garber, NASA History Web Curator
    For further information E-mail histinfo@hq.nasa.gov