APOLLO 13 MISSION REVIEW 

HEAR& 

BEFORE THE 

COMMITTEE ON 

AERONAUTICAL AND SPACE SCIENCES 

UNITED STATES SENATE 

NINETY-FIRST CONGRESS 

SECOR’D SESSION 

47476 0 

JUR’E 30, 1970 

Printed for the use of the 

Committee on Aeronautical and Space Sciences 

U.S. GOVERNMENT PRINTING OFFICE 

WASHINGTON : 1970 
COMMITTEE ON AEROKAUTICAL AND SPACE SCIENCES 

CLINTON P. ANDERSON, New Mexico, Chairman 

RICHARD B. RUSSELL, Georgia 

WARREN G. MAGNUSON, Washington 

STUART SYMINGTON, bfissouri 

JOHN STENNIS, Mississippi 

STEPHEN M. YOUNG, Ohio 

THOJfAS J. DODD, Connecticut 

HOWARD W. CANNON, Nevada 

SPESSARD L. HOLLAND, Florida 

MARGARET CHASE SMITH, Maine 

CARL T. CURTIS, Nebraska 

MARK 0. HATFIELD, Oregon 

BARRY GOLDWATER, Arizona 

WILLIAM B. SAXBE, Ohio 

RALPH T. SMITH, Illinois 

J4MES J. GEHRIG, Stad Director 

EVERARD H. SMITH, Jr., Professional staff Member 

Dr. GLEN P. WILSOS, Professional #tad Member 

CRAIG VOORHEES, Professional Staff Nember 

WILLIAM PARKER, Professional Staff Member 

SAM BOUCHARD, Assistant Chief Clerk 

DONALD H. BRESNAS, Research Assistant 

(11) 
Page 

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CONTENTS 

Tuesday, June 30, 1970 : Opening statement by the chairman, Senator Clinton P. Anderson-__- 

Review Board Findings, Determinations and Recommendations-----_ 

Testimony ofDr. Thomas 0. Paine, Administrator of NASA, accompanied by 

Edgar M. Cortright, Director, Langley Research Center and 

Chairman of the dpollo 13 Review Board ; Dr. Charles D. Harrington, Chairman, Aerospace Safety Advisory Panel ; Dr. 

Dale D. Myers, Associate Administrator for Manned Space 

Flight, and Dr. Rocco A. Petrone, hpollo Director -___________ 21, 30 

Edgar 11. Cortright, Chairman, hpollo 13 Review Board-------- 21,27 

Dr. Dale D. Mvers. Associate Administrator for Manned SDace 

LIST OF ILLUSTRATIOSS 

1. Internal coinponents of oxygen tank So. 2 ---_____-_________________ 2. Oxygen tank wiring and lines _-_-_______________-_________________ 

3. Photograph of switch test -_-_______---____-______________________- 4. Photograph of wiring test ___________--_-_________________________- 

5. Review Board organization chart __-____--_________-_______________ 6. Apollo/Saturn space vehicle---_---_--__-_-----__-_----__----___--- 

7. Service module--_-_---____-__----_-___----_-_-_-_--_--_----__-_-_ 

8. Arrangement of fuel cells and cryogenic in Bay 4 _______________-___ 

9. Photograph of Bay 4 showing major elements _____________-______--- 10. Photograph of Bay 4 showing major eleineiits _________-____-_______ 11. Photograph of Bay 4 showing major elements _______________________ 12. Oxygen tank shell__-__-__-----_--__-__------___------_--_-__--_-- 

13. Routing of Tvires 

14. Crew photograph shows missing panel _____________-_______________ 16. Accident eventu____-_--___-___--__-_-__-___---___-______--___--__- 

16. Flight plan after accident_-_-____-_--_-___--_-_-_____-_____---__-- 

17. Flight plan after accident _____________-___--_____________________- 

18. Thermostatic snitch __--_______________________-_-_____-____

19. Condition of wires _____________________________--___-________--___ 

20. Condition of wires _____-___________________________________-_____- 

21. Review Board Recommendation KO, 1 ________________-___-__-__---- 22. Redesign approach to oxygen system ______--_______--______________ 23. Modified oxygen storage system _____________-______________________ 24. Review Board Recommendation So. 2 ______________________________ 25. Review Board Reconimendation So. 3h _____________________________ 26. Review Board Recommendation KO. 3B ___-______-____________--_-_ 27. Review Board Recoininendation So. 3C ____________________________ 28. Review Board Recommendation So. 3D ____-_____________-_________ 

29. Review Board Recomniendatioii So. 4 _____________________________ 30. Review Board Recommendation So. 3 ________-_____________________ 31. Review Board Recomn~endation So. 6 __-___________-_____--____--- 

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6' 
APOLLO 13 MISSION REVIEW 

TUESDAY, JUNE 30, 1970 

UNITED STATEE SENATE, 

COMMITTEE ON AERONAUTICAL 

AND SPACE SCIENCES, 

Washington. D.C. 

The committee met, pursuant to call, at 10:05 a.m., in room 235, 

Old Senate Office Building, Senator Clinton P. Anderson (chairman) 

presiding. 

Present : Senators Anderson, Stennis, Holland, Curtis, Goldwater, 

and Smith of Illinois. 

*4lso present: James J. Gehrig, staff director; Everard H. Smith, 

Jr., Dr. Glen P. Wilson, Craig Voorhees, and William Parker, professional staff members ; Sam Bouchard, assistant chief clerk ; Donald 

H. Brennan, research assistant, and Mary Rita Robbins, clerical assistant. 

OPENING STATEMENT BY THE CHAIRMAN 

The CHAIRMAS. The hearing \Till come to order. 

Today the committee will continue its review of the Apollo 13 mission initiated on April 24,1970, and will hear testimony on the findings 

of the Apollo 13 Revien- Board and the status of the Apollo program. 

Dr. Thomas 0. Paine, Administrator of NASA, accompanied by 

NASA manned space flight officials. Dr. Dde D. Myers, Associate 

&4dministrator for Manned Space Flight, and Dr. Rocco Petrone, 

Apollo Program Director, will testify on NAS14 actions to be taken 

with respect to the findings and recommendations of the Apollo 13 

RevieF Board, on NASA plans for modifying the Apollo hardware 

and 011 plans for resumption of the L4pollo flight program. 

Mr. Edgar M. Cortright, Director of NASA's Langley Research 

Center, who is Chairman of the Apollo 13 Review Board, will present 

a brief summary highlighting the Revien- Board activities and provide any additional information that may have developed from tests 

in process when the Board report was released. 

Also accompanying Dr. Paine today is Dr. Charles D. Harrington, 

Chairman of the XASA Aerospace Safety Advisory Panel. Dr. Paine 

rill submit for the record a copy of a letter he received from Chairman Harrington giving the Panel's comment on the procedures and 

findings of the Apollo 13 Review Board. Dr. Harrington will not 

present a statement but is ]>resent to respond to questions committee 

members may have regirding the Panel's oversight of the activity 

of the Review Board. 

If there is no objection, I will have chapter 5 of the report of the 

Apollo 13 Revien- Board placed at an appropriate point in the record. 

(1) 
2 

This chapter contains the findings, determinations, and recommendations of the Board resulting from several nTeeks of intensive investigation into the cause of the failure n-hich necessitated aborting the Apollo 

13 mission. 

The CHAIRJIAN. Senator Smith is not here at the moment. Go ahead, 

Dr. Paine. 

(Chapter 5 referred to above is as follows :) 

CHBPTER GFISDISGS, DETERMISATIONS, ASD RECObIJIESDATIONS 

PART 1. ISTRODUOTIOX 

The following findings, determinations, and recommendations are the product 

of about 7 weeks of concentrated review of the Apollo 13 accident by the hpollo 

13 Review Board. They are based on that review, on the accident investigation by 

the Manned Spacecraft Center (JISC) and its contractors. and on an extensive 

series of special tests and analyses performed by or for the Board and its Panels. 

Sufficient work has been done to identify and understand the nature of the 

malfunction and the direction which the corrective actions must take. All indicabions are that an electrically initiated fire in oxygen tank no. 2 in the service 

module ( SJI) was the cause of the accident. Accordingly, the Board has concen- trated on this tank; on its design, manufacture, test. handling. checkout, use, 

failure mode. and eventual effects on the rest of the spacecraft. The accident is 

generally understood, and the most probable cause has been identified. However, 

at the time of this report, some details of the accident are not completely clear. 

Further tests and analyses, which will be carried out under the overall direction 

of MSC, will continue to generate new information relatire to this accident. It is 

possible that this evidence may lead to conclusions differing in detail from those 

which can be drawn now. However, it is most unlikely Ohat fundamentally different results will be obtained. 

Recommendations are provided as to the general direction \vhich the correctire 

actions should take. Significant modifications should be made to the SJI oxygen 

storage tanks and related equipments. The modified hardware should go through a rigorous requalification test program. This is the responsibility of the Apollo 

organization in the months ahead. 

In reaching its findings, determinations, and recommendations, it was neces- sary for the Board to review critically the equipment and the organizational 

elements responsible for it. It was found that the accident was not the result 

of a chance malfunction in a statistical sense, but rather resulted from an iinusual combination of mistakes, coupled with a somewhat deficient and unforgiving design. In brief, this is what happened : 

(a) After assembly and acceptance testing, the oxygen tank So. 2 which flew 

on Bpollo 13 was shipped from Beech Aircraft Corporation to Sorth American 

Rockwell (KR) in apparently satisfadory condition. 

(b) It is now known, however, that the tank contained two protective thermostatic switches on the heater assembly, which n-ere inadequate and would subsequently fail during ground test operations at Kennedy Space Center (KSC). 

(c) In addition, it is probable that the tank 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. (a) In itself, the displaced fill tube assembly was not particularly serious. 

but it led to the use of improvised detanking Irocedures at KSC which almost 

certainly set the stage for the accident. 

(e) Although Beech did not encounter any problem in detanking during acceptance tests, it was not possible to detank oxygen tank So. 2 using normal 

procedures at KSC. Tests and analyses inclicate that this was due to gas leakage 

through the displaced fill tube assembly. 

(f) The special detanking procedures at KSC subjected to 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 7-iolate the specifications 

which governed the operation of the heaters at KSC. 
3 

(9) In reviewing these procedures before the flight, officials of XASA, NR, 

and Beech did not recognize the possibility of damage due to overheating. Many 

of these officials were not aware of the extended heater operation. In any event, 

adequate thermostatic switches might have been expected to protect the tank. 

(11) A number of factors contributed to the presence of inadequate thermostatic switches in the heater assembly. The original 1962 specifications from 

SR to Beech Aircraft Corporation for the tank and heater assembly specified 

the use of 28 V dc power, which is used in the spacecraft. In 1965, SR issued a revised specification which stated that the heaters should use a 65 V dc power 

supply for tank pressurization; this was the power supply used at KSC to reduce pressurization time. Beech ordered switches for the Block I1 tanks but 

did not change the switch specifications to be compatible with 65 V dc. 

(i) The thermostatic switch discrepancy was not detected by SASA, XR, 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. 

(j) The thermostatic switches could accommodate the 65 V dc during tank 

pressurization because they normally remained cool and closed. However, they 

could not open without damage with 65 T' dc power applied. They were never required to do so 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 inoperatire as protectire thermostats. 

(k) Failure of the therniostatic switches to open could have been detected 

at RSC if switch operation had been checked by obserring heater current readings on the oxygen tank heater control panel. Although it was not recognized 

at that time, the tank temperature readings indicated that the heaters had 

reached their temperature limit and switch opening should have been expected. 

(1) As shown by subsequent tests, failure of the thermostatic switches probably 

permitted the temperature of the heater tube assembly to reach about 1000" F 

in spots during the continuous %hour period of heater operation. Such heating 

has been shown by tests to severely damage the Teflon insulation on the fan 

motor mires in the vicinity of 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. 

(m) It was not until nearly 56 hours into the mission, however, that the 

fan motor miring, possibly moved by the fan stirring, short circuited and ignited 

its insulation by means of an electric arc. The resulting combiistion in the oxygen 

tank probably overheated and failed the wiring conduit where it enters the tank, 

and possibly a portion of the tank itself. 

(n) 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. 

The accident is judged to have been nearly catastrophic. Only outstanding 

performance on the part of the crew, Mission Control, and other members of the 

team which supported the operations successfully returned the crew to Earth. 

In investigating the accident to Apollo 33. the Board has also attempted to 

identify those additional technical and mangement lessons which can be applied 

to help assure the success of future space flight missions ; several recommendations of this nature are included. 

The Board recognizes that the contents of its report are largely of a critical 

nature. The report highlights in detail faults or deficiencies in equipment and 

procedures that the Board has identified. This is the nature of a review board 

report. 

It is important, however, to view the criticisms in this report in a broader 

contest. The Apollo spacecraft system is not without shortcomings, but it is the 

only srsteni of its type erer built and successfnlly demonstrated. It has flown 

to the Moon five times and landed twice. The tank which failed, the design of 

which is criticized in this report, is one of a series which had thousands of hours 

of successful operation in space prior to Apollo 13. 

While the team of designers, engineers, and technicians that build and 

operate the Apollo spacecraft also has shortcomings, the accomplishments speak 

for themselves. By hardheaded self-criticism and continued dedication, this team can maintain this nation's preeminence in space. 
4 

PART 2. ASSESSMEWT OF ACCIDENT 

FAILURE OF OXYGEN TANK NO. 2 

1. Findings 

(a) The A%poll~ 13 mission was aborted as the direct result of the rapid loss 

of oxygen from oxygen tank NO. 2 in the SJI, followed by a gradual loss of 

oxygen from tank no. 1, and a resulting loss of power from the oxygen-fed fuel 

cells. 

(b) There is no evidence of any forces external to oxygen tank no. 2 during the flight which might have caused its failure. 

(c) Oxygen tank no. 2 contained materials, including Teflon and aluminum, 

which if ignited mill burn in supercritical oxygen. 

(d) Oxygen tank no. 2 contained potential ignition sources : electrical miring, unsealed electric motors, and rotating aluminum fans. 

(e) During the special detanking of oxygen tank no. 2 following the countdown demonstration test (CDDT) at KSC, the thermostatic switches on the 

heaters were required to open while powered by 65 l’ dc in order to protect the 

heaters from overheating. The switches were only rated at 30 l’ dc and hare 

been shown to weld closed at the higher voltage. 

(f) Data indicate that in flight the tank heaters located in oxygen tanh no. 1 and no. 2 operated normally prior to the accident, and they were not on at the 

time of the accident. 

(9) The electrical circuit for the quantity probe would generate only about 

7 millijoules in the event of a short circuit and the temperature Sensor wires less 

than 3 millijouleis per second. 

(h) Telemetry data immediately prior to the accident indicate electrical disturbances of a character which would be caused by short circuits acconipanied by 

electrical arcs in the fan motor or its leads in oxygen tank no. 2. 

(4) The pressure and temperature within oxygen tank no. 2 rose abnormally 

during the 1% minutes immediately prior to the accident. 

Determinations 

(1) The cause of the failure of oxygen tank no. 2 was cornbustion within the 

tank. 

(2) Analysis showed that the electrical energy flowing into the tank could not 

account for the observed increases in pressure and temperature. 

(3) The heater, temperature sensor, and quantity probe did not initiate the 

accident sequence. 

(4) The cause of the combustion was most probably the ignition of Teflon wire 

insulation on the fan motor wires. caused by electric arcs in this wiring. 

(5) The protective thermostatic switches on the heaters in oxygen tank no. 2 

failed closed during the initial portion of the first special cletanking operation. 

This subjected the wiring in the vicinity of thp heaters to very high temperatures \\-hich have been subsequently shown to sererely degrade Teflon insulation. 

(6) The telenietered data indicated electrical arcs of sufficient energy to ignite the Teflon insulation, as verified by subsequent tests. These tests also rerified 

that the 1-ampere fuses on the fan motons m*ould pass sufficient energy to ignite 

the insulation by the mechanism of an electric arc. The coinbustion of Teflon wire insulation alone could release sufficient 

heat to account for the observed iiicreases in tank pressure and local temperature, and could locally overheat and fail the tank or its associated tubing. The 

possibility of such failure at the top of the tank was demonstrated by <.ubsequent 

tests. 

(8) The rate of flame propagation along Teflon-insulated wires as meamred 

in subsequent tests is consistent with the indicated rates of pressure rise within 

the tank. 

2. Findings 

(7) 

SECONDARY EFFECTS OF TASK FAILCRE 

(a) Failure of the tank was acconipanied by several events including : 

A “bang” as heard by the crew. Spacecraft motion as felt by the crelv and as measured by the attitude 

Momentary loss of telemetry. 

Closing of several valves by shock loading. 

control system and the accelerometers in the command module (chi). 
Loss of integrity of the oxygen tank no. 1 system. 

Slight temperature illcreases in bay 4 and adjacent sectors of the SX 

Loss of the panel covering bay 4 of the SlI, as observed and photographed 

Displacement of the fuel cells as photographed by the wen-. Damage to the high-gain antenna as photographed by the crew. (b) The pallel covering of bay 4 could be blown off by pressurization of the 

bay. About 2.5 psi of uniform pressure in bay 4 is required to blow off the panel. 

(c) The rarious bays and sectors of the SM are interconneded with open 

pacsages so that all would be pressurized if any oiie were supplied with a pressurant at a relatively slo~r rate. 

(rl) The CJl attachments would be failed by an average pressure of about 

10 psi on the CJI heat shield and this would separate the CBI from the SR1. 

Deternzinations 

(1) Failure of the oxygen tank no 2 caused a rapid local pressurization of 

bay 4 of the SJI by the high-pressure oxygen that escaped froin the tank. This 

pre+sure pulse may hare blown off the panel covering bay 4. This possibility was substantiated by a series of special tests. 

(2) The pressure pulse from a tank failure might have been augnieiited by 

combustion of Mylar or Kapton insulation or both when subjected to a stream 

of oxygen and hot particles emerging from the top of the tank, as demonstrated 

in subsequent tests. 

(3) Combustion or vaporization of the Mylar or Kapton might account for 

the discoloration of the SJI engine nozzle as observed and photographed by the 

crew. 

(4) Photographs of the SJl by the crew did not establish the condition of the 

oxygen tank no. 2. 

(5) The high-gain antenna damage probably resulted from striking by the 

panel, or a portion thereof, as it left the SlI. 

(6) The loss of pressure on oxygen tank no. 1 and the subsequent loss of 

power resulted from the tank no. 2 failure. 

(7) Telemetry, although good, is insufficient to pin down the exact nature, 

sequence, and location of each event of the accident in detail. 

(8) The telenietv data. crew testiniony, lihotographs, and special tests and 

analyses already completed are sufficient to understand the problem and to proceed with correctire actions. 

by the crew. 

OXYGER- TAiTK XO. 2 DESIGS 

3. Findings 

combustible material, and potential ignition sources. 

handling problems of the ox)-gen supply system. 

and management of the oxygen cupply. 

(a) The cryogenic oxygen storage tanks contained a combination of oxidizer, 

(6) Supercritical oxygen was used to minimize the weight, volume, and fluid- 

(c) The heaters, fans, and tank instrumentrition are used in the measurement 

Determinations 

(1) The storage of supercritical oxygen was appropriate for the Ap0ll0 system. 

(2) Heaters are required to maintain tank pressure as the oxygen supply i. 

used. 

(3) Fans were used to prevent excwsire pressure drops due to stratificatioli. 

to mix the oxygen to iinprore accuracy of quantity measurements, and to insure 

adequate heater input at low densities and high oxygen utilization rates. The 

need for oxygen qtirring on future fliglitf requires further investigation. 

(4) The amount of iiiaterial in the tank n-hich could be ignited ancl burnecl ill 

the giren enrironment coulcl have been reducecl significantly. 

(3) The potential ignition sources conctituted an undue hazard when COILsidered in the light of the particular tanli design with its assembly difficulties. 

(6) SALSA, the iiriiiie contractor. and the \upplier of tlie tank Irere rlot fniiy aware of tlie extent of this hazard 

(7) .Examination of the high-prepsure oxygen system in tlie serrice liloduie 

follolvmg the Apollo 204 fire. Tvhich directed attention to the danger of fire in a pure oxygen environment, failed to recognize the deficiencies of the tanli 
6 

PREFLIGHT DAMAGE TO TANK WIRIXG 

4. Findings 

(a) The oxygen tank no. 2 heater assembly contained two thermostatic switches 

designed to protect the heaters from overheating. 

(b) The thermostatic switches were designed to open and interrupt the heater 

current at 80"*10" F. 

(0) The heaters are operated on 28 V dc in flight and at NR. 

(d) The heaters are operated on 65 V ac at Beech Aircraft Corporation and 

65 V dc at the Kennedy Space Center. These higher voltages are used to accelerate 

tank pressurization. 

(e) The thermostatic switches were rated at 7 amps at 30 V dc. While they would carry this current at 65 V dc in a closed position, they would fail if they 

started to open to interrupt this load. 

(7) Neither qualification nor acceptance testing of the heater assemblies o> 

the tank8 required thermostatic switch opening to be checked at G T' dc. The 

only test of switch opening vas a continuity check at Beech in which the switch was cycled open and closed in an oven. (g) The thermostatic switches had never operated in flight because tMs muld 

only happen if the oxygen supply in a tank were depleted to nearly zero. (Ii ) The thermostatic switches had never opepated on the ground under load 

because the heaters had only been used mith a relatively full tank vhich kept the switches cool and closed. 

(i) During the CDDT, the oxygen tank no. 2 would not detank in a normal manner. On March 27 and 28, a special detanking procedure was followed which 

subjected the heater to about 8 hours of continuous openation until the tanks mere nearly depleted of oxygen. 

(j) X second special detanking of shorter duration followed on March 30, 

1970. 

(k) The oxygen tanks had not been qualification tested for the conditions 

encountered in this procedure. However, specified allowable heater voltages and 

currents were not exceeded. 

(1) The recorded internal tank temperature went off-scale high early in the 

special detanking. The therniozitatic switches would normally open at this point 

but the electrical records shorn no thermostatic switch operation. These indications n-ere not detected at the time, 

(m) The oxygen tank heater controls at KSC contained ammeters which 

would have indicated thermostatic switch operation. 

D et evminat ions 

(1) During the special detanking of March 27 and 28 at KSC, when the heater7 

in oxygen tank No. 2 were left on for an extended period, the thermostatic 

sdtches started to open mhile powered by 65 Tr dc and were probably welded 

shut 

(2) Failure of the thermostatic sn'itches to open could hare been detected at 

KSC if switch operation had been checked by observing heater current readings on the oxygen tank heater control panel. Although it mas not recognized at the 

time, the tank temperature readings indicated that the heaters had reached their 

temperature limit and switch opening should hare been expected. 

(3) The fact that the switches were not rated to open at 65 V dc was not 

detected by NASA, R'R, or Beech in their reviews of documentation or in qualification and acceptance testing. 

(4) The failed switches resulted in severe overheating. Subsequent test.; 

showed that heater assembly temperatures could hare reached about 1000" F. 

(.?) The high temperatures severely damaged the Teflon insulation on thr 

wiring in the vicinity of the heater assembly and set the shge for subsequent 

short circuiting. As shown in subsequent tests, this d'amage could range from cracking to total oxidation and disappearance of the insulation. 

(6) Durinq and following the special detanking, the oxygen bank no. 2 vas in a hazardous condition wheneuer it contained oxygen and was electrically 

energized. PART 3. SUPPORTISG COSRIDERATIOSS 

DESIGS, MANUFACTURING, ASD TEST 

5. Finding 

718, and is moderately stressed at normal operating pres711re 

The preesure vessel of the supercritical oxygen tank is conqtructed of Inconel 
7 

Determination 

From a structural viewpoint, the supercritical oxygen pressure vessel is 

quite adequately designed, employing a tough material well chosen for this 

application. The stress analysis and the results of the qualification burst 

test program confirm the ability of the tank to exhibit adequate performance 

in its intended application. 

6. Findings 

(a) The oxygen tank design includes two unsealed electric fan motors imxersed in supercritical oxygen. 

(b) Fan motor? of this design have a test history of failure duriiig Receptance test which includes phase-to-phase and phase-to-ground faults. 

(c) The fan motor stator windings are constructed with Teflon-coated, ce- ramic-insulated, iiuinber 36 AWG wire. Full phase-to-phase and phase-to-ground 

insulation is not used in the motor design. 

(d) The motor case is largely aluminum. 

(1) The stator Ivinding insulation is brittle and easily fractured during mai1l1- 

facture of the stator coils. 

(2) The use of these motors in supercritical oxygen was a questionable 

practice. 

7. Findings 

(a) The cryogenic oxygen storage tanks contained materials that could be 

ignited and which will burn under tlie conditions prevailing within the tanli, 

including Teflon, aluminum, solder, and Drilube 822. 

( b) The tank contained electrical wiring exposed to the supercritical oxygen. 

The wiring was insulated with Teflon. . (0) Some miring was in close proximity to heater elements and to the ro- tating fan. 

(d) The design was such that the assembly of the equipment was essentially 

“blind” and not amenable to inspection after conipletion. 

(e) Teflon insulation of the electrical wiring inside the cryogenic oxygen 

storage tanks of the SJI was exposed to relatively sharp metal edges of tank 

inner parts during manufacturing assembly operations. 

(f) Portions of this n-iring remained unsupported in the tank on completion 

of assembly. 

Determinations 

(1) The tank contained a hazardous combination of materials and potential ignition sources. 

(2) Scraping of the electrical wiring insulation against metal inner parts of 

the tank constituted a substantial cumulative hazard during assembly. handling. 

test. checkout. and operational use. 

(3) “Cold flow” of the Teflon insulation, when pressed against metal cornerh 

within the tank for an extended period of time, could result in an eventual 

degradation of iiisulation protection. 

(4) The externally applied electrical tests (600-volt Hi-pot) could not reveal 

the extent of such possible insulation daniagr but could only indicate that the 

relative positions of the wires at the time of the tests were such that tlie separation or insulation would rrithstand the 500-volt potential without electrical 

breakdorm. 

(5) The design was such that it was clifficult to insure against these hazards 

(6) There is no evidence that the n-iring was damaged during manufacturing. 

9. Findings 

(a) Dimensioning of the short Teflon and Inconel tube segments of the cryo- genic oxygen storage tank fill line vas such that looseness to the point of inconiplete connection mas possible in the event of worst-case tolerance buildup. 

(b) The insertion of these segments into the top of the tank quantity probe assembly at the point of its final closure and welding was difficult to achieve. 

(c) Probing with a hand tool was used in manufacturing to compensate for 

limited risibility of the tube segment positions. 

Determination 

Determinations 

It was possible for a tank to have been assembled with a set of relatively loose 

fill tube parts that could go undetected in final inspection and be subsequently 

displaced. 
8 

10. Fhdings 

(a) The Apollo spacecraft systeni contains numerous pressure vessels, nlany 

of which carry oxidants, plus related ralves and other plumbing. 

(b) Investigation of potential hazards associated with these other systems was not complete at the time of the report, but is being pursued by the i\Lannd 

Spacecraft Center. 

(c) One piece of equipiiient. the fuel cell oxygen supply ralre module. has 

been identified as containing a similar combination of high-pressure oxygen. 

Teflon. and electrical wiring as in the oxygen tank no. 2. The wiring is unfused 

and is routed through a 10-amp circuit breaker. 

Determination 

The fuel cell oxygen supply ralve module has been identified as potentially 

hazardous. 

11. Findings 

(a) In the normal sequence of cryogenic oxygen storage tank integration alld 

checkout, each tank undergoes shipping. assembly into an oxygen shelf for a service module, factory transportation to facilitate shelf assembly test, and then 

integration of shelf assembly to the SJI. 

(b) The SJI undergoes factory transportation, air shipment to KSC. and 

subsequent ground transportation and handling. 

Determination 

There n7ere environments during the normal sequence of operations subsequent to the final acceptance tests at Beech that could cause a loose-fitting set 

of fill tube parts to become displaced. 

12. Findings 

(a) At North American Rockwell, Downey, California, in the attempt to 

reinove the oxygen shelf assembly from ShI 106, a bolt restraining the inner 

edge of the shelf was not removed. 

(b) Attempts to lift the shelf with the bolt in place broke the lifting fixture, 

thereby jarring the oxygen tanks and \-alves. 

(c) The oxygen shelf assembly incorporating S/N XTA0008 in the tank no. 2 

position, which had been shaken during removal froin SM 106, was installed in 

SlI 109 one month later. 

(d) An analysis, shelf inspection, and a partial retest emphasizing electrical 

continuity of internal wiring were accomplished before reinstallation. 

Determ‘nations 

(1) Displacement of fill tube parts could have occurred, during the “shelf 

(2) Other damage to the tank may have occurred from the jolt, but special 

(3) The “shelf drop” incident was not brought to the attention of project 

drop” incident at the prime contractor’s plant, without detection. 

tests and analyses indicate that this is unlikely. 

officials during subsequent detanking difficulties at KSC. 

13. Binding 

Detanking, expulsion of liquid oxygen out the fill line of the oxygen tank by warm gas pressure applied through the rent line, was a regular activity at Beech 

Aircraft, Boulder, Colorado, in emptying a portion of the oxygen used in end-item 

acceptance tests. 

Determination 

The latter stages of the detanking operation on oxygen tank no. 2 conducted 

at Beecll on February 3, 1967, were similar to the standard procedure followed 

at KSC during the CDDT. 

14. Findings 

(a) The attempt to detnnk the cryogenic oxygen tanks at KSC after the CDDT 

by the standard procedures on March 23. 1970, was unsuccessful with regard to 

tank no. 2. 

(b) A special detanking procedure was used to empty oxygen tank no. 2 after 

CDDT. This procedure involved continuous protracted heating with repeated 

cycles of pressurization to about 300 psi with warm gas followed by venting. 

(c) It was employed both after CDDT and after a special test to verify that 

the tank could be filled. 
9 

(d) There is no indication from the heater voltage recording that the thermostatic switches functioned and cycled the heaters off and on during these special 

detanking procedures. (e) At the conipletion of detaiiking following CDDT, the switches are only 

checked to see that they remain closed at -75" F as the tank is warmed up. They are not checked to verify that they will open at $80" F. 

(f) Tests subsequent to the flight shon-ed that the current associated mith the 

KSC 65 V dc ground po\vering of the heaters would cause the thermostatic switch 

contacts to weld closed if they attempted to interrupt this current. 

(9) d second test showed that without functioning thermostatic switches, 

temperatures in the 800" to 1000' F range would exist at locations on the heater 

tube assembly that were in close proximity with the motor Fires. These temperatures are high enough to damage Teflon and inelt solder. 

Determimations 

(1) Oxygen tank no. 2 (XTA 0008) did not detank after CDDT in a manlier 

comparable to its performance the last time it had contained liquid oxygen, i.e., 

in acceptance test at Beech. 

(2) Such evidence indicates that the tank had undergone smie change of 

internal configuration during the intervening events of the previous 3 years. 

(3) The tank conditions during the special detanking procedures were outside 

all prior testing of Apollo CSBI cryogenic oxygen storage tanks. Heater assembly 

temperatures measured in subsequent tests exceeded 1000" I?. 

(4) Severe damage to the insulation of electrical wiring internal to the tank. as determined from subseqnent tests, resultecl from the special procedure. 

(5) Damage to the insulation, particularly on the long unsupported lengths of 

wiring, may also hare occurred due to boiling associated with this procedure. 

(6) JISC, KSC, and SR personnel did not know that the thermostatic switches were not rated to open with 65 V dc GSE power applied. 

15. Findings 

(a) The change in detanking procedures on the cryogenic oxygen tank was made in accordance with the existing change control system during final launch 

preparations for dpollo 13. 

(b) Launch operations personnel who made the change did not have a detailed 

understanding of the tank internal components, or the tank history. They made 

appropriate contacts before iiiaking the change. 

(c) Communications, primarily by telephone, among JISC, KSC, SR, and 

Beech personnel during final launch preparations regarding the cryogenic oxygen 

system included incomplete and inaccurate information. 

(d) The JISC Test Specification Criteria Document (TSCD) which was used 

by KSC in preparing detailed tank test procedures states the tank allowable 

heater voltage and current as 63 to 8.5 T' clc and 9 to 17 amperes with no restrictions on time. 

D eterni ina t ions 

(1) SR and hISC personnel who prepared the TSCD did iiot know that the 

(2) Launch operations personnel assumed the tank was protected from over- 

(3) Launch operations personnel at KSC stayed within the specified tank 

tank heater tlieniostatic snitches would not protect the tank. 

heating by the switches. 

heater voltage and current liniits during the detanking at KSC. 

16. Findings 

(a) After receipt of the Block I1 oxygen tank specifications fro111 SR. whicli 

required the tank heater assembly to operate with 65 T' dc GSE power only during 

tank Iressurization, Beech Aircraft did not require their Block I thermostatic 

switch supplier to make a change in tlie switch to operate at the higher voltage 

(b) SR did not reT-iew the tank or heater to assure compatibility between 

tlir snitch and the GSE. 

ic) JlRC did not review the tank or heater to ns\ure compatibility hetween the switch and tlie GSE. 

(d) No tests were specified by JISC. SR. or Reecli to check this slyitch uncler 

load. 

Dctcrniinations 

(1) SR and Beech specifications governing the powering and the therniostatio 

switch protection of the heater assemblies were inadequate. 
10 

(2) The speciflcations governing the testing of the heater assemblies were inadequate. 

17. Finding 

given consideration in the decision to fly oxygen tank no. 2. 

The hazard associated with the long heater cycle during detanking was not 

Determinations 

(1) MSC, KSC, and NR personnel did not know that the tank heater thermo- 

(2) If the long period of continuous heater operation with failed thermostatic 

static switches did not protect the tank from overheating. 

switches had been known, the tank would have been replaced. 

18. Findings 

(5) Management controls requiring detailed reviews and approvals of design, manufacturing processes, assembly procedures, test procedures, hardware 

acceptance, safety, reliability, and flight readiness are in effect for all Apollo 

hardware and operations. 

(b) When the Apollo 13 cryogenic ‘oxygen system was originally designed, the 

management controls were not defined in as great detail as they are now. 

Determination 

From review of documents and interviews, it appears that the management 

controls existing at that time were adhered to in the case of the cryogenic oxygen 

system incorporated in Apoll~ 13. 

19. Finding 

leak through the vent quick disconnect, which was corrected. 

The only oxygen tank no. 2 anomaly during the final countdown was a small 

Deterrnilzation 

KO indications of a potential inflight malfunction of the oxygen tank no. 2 

were present during the launch countdown. 

20. Findings 

(5) The center engine of the S-I1 stage of the Saturn V launch vehicle prematurely shut down at 132 seconds due to large 16 hertz oscillations in thrust 

chamber pressure. 

(b) Data indicated less than 0.lg vibration in the (YM. 

Determinationa 

(1) Investigation of this S-I1 ,anlomaly was not within the purview of the 

(2) The resulting oscillations or vibration of the space vehicle probably did 

MISSION EVENTS THROUGH ACCIDENT 

Board except insofar as it relates to the Apollo 13 accident. 

not affect the oxygen tank. 

21. Findings 

(5) Fuel cell current increased between 46 :40 :05 and 46 :40 :08 indicating that the oxygen tank no. 1 and tank no. 2 fans were turned on during this interval. 

(b) The oxygen tank no. 2 quantity indicated off-scale high at 46:40:08. 

(1) The oxygen tank no. 2 quantity probe short circuited at 46 :40 :08. 

(2) The short circuit could have been caused by either a completely loose fill 

tube part or a solder splash being carried by the moving fluid into contact 

with both elements of the probe capacitor. 

22. Findings 

fans at 55 :53 :06. 

Deteminations 

(5) The crew acknlowledged Mission Control’s request to turn on the tank 

(b) Spacecraft current increased by 1 ampere at 55 :53 :IS. 

(c) The oxygen tank no. 1 pressure decreased 8 psi at 55 :63 :19 due to normal 

destratification. 

Determination 

The fans in oxygen tank no. 1 were turned on and began rotating at 55 :53 :19. 
11 

2.3. Findings 

(a) Spacecraft current increased by 1% amperes and ac bus 2 voltage 

decreased 0.6 volt at 55 :53 :20. 

(a) Stabilization and Control System (SCS) gimbal command telemetry 

channels, \I-hich are sensitive indicators of electrical transients associated with 

sn*itching on or off of certain spacecraft electrical loads, shoved a negative initial transient during oxygen tank no. 2 fan turnon cycles and a positive 

initial transient during oxygen tank no. 2 fan turnoff cycles during the Apollo 13 

mission. d negative initial transient was measured in the SCS at 55 :53 :20. 

(c) The oxygen tank no. 2 pressure decreased about 4 psi when the fans were turned on at 55 :53 :21. 

Determinations 

(1) The fans in oxygen tank no. 2 were turned on at 55 :53 :20. 

(2) It csnnot be determined whether or not they were rotating because the 

pressure decrease was too small to conclusively show destratification. It is 

likely that they were. 

24. Fiiiding 

were measured in lac bus 2 at 55 :53 :23. 

An 11.1-amp spike in fuel cell 3 current and a monientary 1.2-volt decrease 

Determinations 

(1) A short circuit occurred in the circuits of the fans in oxygen tank no. 2 

which resulted in either blown fuses or opened wiring, and one fan ceased to 

function. 

(2) The short circuit probably dissipated an energy in excess of 10 joules 

which, as shown in subsequent tests, is more than sufficient to ignite Teflon 

wire insulation by means of an electric arc. 

25. Findings 

(a) d momentary ll-volt decrease in ac bus 2 voltage was measured at 

55 :53 :38. 

(6) A 22.9-amp spike in fuel cell 3 current mas measured at 55 :53 :41. 

(c) After the electrical transients, CM current and ac bus 2 voltage returned 

to the values indicated prior to the turnon of the fans in oxygen tank no. 2. 

Determination 

Two short circuits occurred in the oxygen tank no. 2 fan circuits betffeen 

3.5 :53 :38 and 55 :53 :41 which resulted in either blown fuses or opened wiring, and the second fan ceased to function. 

26. Finding 

Oxygen tank no. 2 telemetry showed a pressure rise from 887 to 954 psia 

between 55 53 :36 and 55 :54 :OO. It then remained nearly constant for about 

15 seconds and then rose again from 954 to 1008 psia, beginning at 55 :54 :16 and 

ending at 55 :54 :45. 

Determinations 

(1) An abnormal pressure rise occurred in oxygen tank no. 2. 

(2) Since no other known energy source in the tank could produce this 

pressure buildup, it is concluded to have resulted from combustion initiated by the first short circuit which started a wire insulation fire in the tank. 

$7. Findings 

psi. 

55 : 54 : 45 to 996 psia at 55 : 54 : 53, at rhich time telemetry data were lost. 

(a) The pressure relief valve was designed to be fully open at about 1000 

(a) Oxygen tank no. 2 telemetry showed a pressure drop from 1008 psia at 

Determination 

verified in subsequent tests. 

28. Findings 

(a) At 55 : 54 : 29, when the pressure in oxygen tank no. 2 exceeded the master 

caution and warning trip level of 975 psia, the CJl master alarm was inhibited 

by the fact that a warning of low hydrogen pressure was already in effect, and 

neither the crew nor Mission Control mas alerted to the pressure rise. 

This drop resulted from the normal operation of the pressure relief valve as 
12 

(6) The master caution and warning system logic for the cryogenic system is 

such that an out-of-tolerance condition of one measurement which triggers a master alarm prevents another inaster alarm from being generated when any 

other parameter in the same system becomes out-of-tolerance. 

(c) The low-pressure trip lerel of the master caution and warning system for the cryogenic storage system is only 1 psi below the specified lower limit of 

the pressure sivitch which controls the tank heaters. A small imbalance in hydrogen tank pressures or a shift in transducer or switch calibration can cause the 

inaster caution and warning to be triggered preceding each heater cycle. This 

occurred several times on Apollo 13. 

(d) A limit sense light indicating abnormal oxygen tank no. 2 pressure should 

have come on in Mission Control about 30 seconds before oxygen tank no. 2 

failed. There is no may to ascertain that the light did, in fact, come on. If it 

did come on, Nission Control did not observe it. 

Deternzinations 

(1) If the pressure switch setting and master caution and warning trip levels were separated by a greater pressure differential, there would be less likelihood 

of unnecessary master alarms. 

(2) With the present master caution and warning system, a spacecraft prob- lem can go unnoticed because of the presence of a previous out-of-tolerance con- dition in the same subsystem. 

(3) Although a master alarm at 55: 54: 29 or observance of a limit sense light in Mission Control could have alerted the crew or Mission Control in sufficient 

time to detect the pressure rise in oxk'gen tank no. 2, no action could have been 

taken at that time to prerent the tank failure. However, the information could 

hare been helpful to Mission Control and the crew in diagnosis of spacecraft 

malfunctions. 

(4) The limit sense system in Mission Control can be modified to constitute a more positive backup warning system. 

29. Finding 

Oxvgen tank no. 2 telemetry showed a temperature rise of 38" F beginning 

at 55 : 54 : 31 sensed by a single sensor which measured local temperature. This 

sensor indicated off-scale lorn at 55 : 54 : 53. 

Deternzinations 

(1) An abnormal and sudden temperature rise occurred in oxygen tank no. 2 at approximately 5.5 : 54: 31. 

(2) The temperature was a local value which rose when combustion had 

proarewed to the ricinity of the sensor. 

(3) The temperature sensor failed at 55 : 54 : 53. 

30. Finding 

Oxygen tank no. 2 telemetry indicated the following changes: (1) quantity 

decreased from off-scale high to off-scale low in 2 seconds at 55 : 54 : 30, (2) quantity increased to 76.3 percent at 55: 54: 32, and (3) quantity vas off-scale high at 

56 : 54 : 51 and later became erratic. 

Deternzinations 

(1) Oxygen tank no. 2 quantity data between 55: 54: 32 and 55: 54: 50 may 

(2) Immediately preceding and following this time period, the inclications were represent ralid measurements. 

caused by electrical faults. 

81. Findings 

(a) At about 65 : 64 : 53, or about half a second before telemetry loss, the Godymounted linear accelerometers in the command module. which are sampled at 

100 times per second, began indicating spacecraft motions. These disturbances were erratic, but reached peak ralues of 1.17.g. 0.6Sg. and 0.6.5g in the X. T. and Z 

directions, respectively, about 13 milliseconds before data loss. 

( b) The body-mounted roll, pitch, and yaw rate gyros showed low-level activity 

for second beginning at 55 : 54 : 53.220. 

(c) The intenratinn accelerometers indicated that a velocitv increment of a])- 

proximately 0.5 fps \vas imparted to the spacecraft betweb 56 : 54 : 53 & 

55 : 54 : 55. 
13 

(d) Doppler tracking data measured an incremental velocity component of 0.26 

fps along a line from the Earth to the spacecraft at approximately 55 : 64 : 55. 

(e j The crew heard a loud “bang” at about this time. 

(f) Telemetry data were lost between approximately 55 : 54 : 53 and 55 : 54 : 56 

and the spacecraft sn-itched from the narrorv-beam antenna to the wide-beam 

antenna. 

and the high-gain antenna to be damaged. 

(g) Crew observations and photographs showed the bay 4 panel to be missing 

Deternaiiz at ions 

(1) The spacecraft was subjected to abnormal forces at approximately 

55 : 54 : 53. These disturbances Jvere reactions rc?sulting froin failure and renting 

of the oxygen tank no. 2 system and subsequeiit separation and ejection of the 

bay 4 panel. 

from bay 4 at the time of panel separation. 

(2) The high-gain antenna was damaged either by the panel or a section thereof 

52. Finding 

cated abnormal increases following reacquisition of data at 5.5 : 54 : 55. 

Temperature sensors in bay 3. bay 4, and the central column of the SJI indiDeternz inat ion 

Heating took place in the SM at approximalely the time of panel separation. 

(a) The telemetered nitrogen pressure in fuel cell 1 was off-scale low at re- 

(21) Fuel cell 1 continued to operate for about 3 minutes past this time. 

(cj The wiring to the nitrogen sensor passes along the top of the shelf which 

33. Finrlitigs 

acquisition of data at 55 : 54 : 55. 

supports the fuel cells immediately above the oxygen tanks. 

Determinations 

(1) The nitrogen pressure sensor in fuel cell 1 or its wiring failed at the 

time of the accident. 

(2) The failure was probably caused by physical damage to the sensor wiring or shock. 

(3 j This is the only knon;n instrumentation eailure outside the oxygen system 

at that time. 

34. Finding 

Oxygen tank no. 1 pressure decreased rapidly froni 879 psia to 782 psia at 

approximately 55 : 54 : 54 and then began to decrease more slowly at 55 : 54 : 56. 

Determination 

A leak caused loss of oxygen from tank no. 1 beginning at approximately 

55 : 54 : 54. 

35. Findings 

(a) Oxygen flow rates to fuel cells 1 and 3 decreased in a 5-second period be. 

ginning at 55 : 54 : 55, but sufficient volume existed in lines feeding the fuel cells 

to allow them to operate about 3 minutes after the oxygen valves were cut off. 

(b) The crew reported at 55: 57: 44 that fire valves in the reaction control 

system (RCS) mere closed. The shock required to close the oxygen supply valves 

is of the same order of magnitude as the shock required to close the RCS 

valves. 

(c) Fuel cells 1 and 3 failed at about 55 : 58. 

Determination 

The oxygen supply valves to fuel cells 1 and 3, and the five RCS valves. were probably closed by the shock of tank failure or panel ejection or both. 

36. Findinys 

(a) Since data presented to flight controllers in Mission Control are updated 

only once per second, the 1.8-second loss of data which occurred in Mission Control vas not directly noticed. However, the Guidance Officer did note and report a “hardware restart” of the Spacecraft computer. This was quickly followed by the 

crew’s report of a problem. 

3fISSIOS ETEXTS AFTER ACCIDEXT 

47-476 0-7&--2 
14 

(a) Immediately after the crew’s report of a “bang” and a main bus B 

undervolt, all fuel cell output currents and all bus voltages were normal, and 

the cryogenic oxygen tank indications were as follows : Oxygen tank no. 1 : Pressure : Severiil hundred psi below normal. QuanOxygen tank no. 2 : Pressure : Off-scale lorn. Quantity : Off-scale high. 

(c) The nitrogen pressure in fuel cell 1 indicated zero, which was incompati- ble with the hydrogen and oxygen pressures in this fuel cell, which were iionnal. 

The nitrogen pressure is used to regulate the oxygen and hydrogen pressure, 

and hydrogen and oxygen pressures in the fuel cell would follow the nitrogen 

pressure. (a) Seither the creir nor Mission Control was aware at the time that oxygei 

tank no. 2 pressure had risen abnormally just before the data loss. 

(c) The flight controllers beliered that a probable cause of these indiclations 

could have been a cryogenic storage system instruinentation failure, and began 

pursuing this line of investigation. 

tity : Kornial. Temperature : Normal. 

Temperature : Off-scale high. 

Deternvination 

Under these conditions it was reasonable to suspect a cryogenic stonage system 

instrumentation problem, and to attempt to yerify the readings before taking 

anF action. The fact that the oxygen tank no. 2 quantity nieasurenient was known 

to hare failed several hours earlier also contributed to the doubt about the creditability of the telemetered data. 

37. Findings 

(a) During the 3 minutes following data loss, neither the flight controllers nor the crew noticed the oxygen flows to fuel cells 1 and 3 mere less than 0.1 lb/hr. 

These were unusually low readings for the current being drawn. 

(6) Fuel cells 1 and 3 failed at about 3 minutes after the datla loss. 

(c) After the fuel cell failures, which resulted in dc main bus B failure and 

the undervoltage condition on dc main bus A, Mission Control diverted its prime 

concern from what mas initially believed to be a cryogenic system instrumenta. 

tion probleni to the electrioal power system. 

(d) Sear-zero oxygen flow to fuel cells 1 ancl 3 was noted after the niain bus R 

failure, but this was consistent with no power output from the fuel cells. 

(e) The flight controllers believed that the fuel cells could hare been disconnected from the busses and directed the crew to connect fuel cell 1 to clc main bus 

h and fuel cell 3 to dc main bus B. 

(f) The crew reported the fuel cells were configured as directed and that the 

talkback indicators confirmed this. 

Deternzilzations 

(1) Vnder these conditions it was logical for the flight controllers to attempt to 

regain power to the busses since the fuel cells might hare been disconnected as a result of a short circuit in the electrical system. Telemetry does not indicate 

whether or not fuel cells are connected to busses, and the available data would not 

distinguish betn-eeii a disconnected fuel cell and a failed one. 

(2) If the crew had been aivare of the reactant rake closure, they could hare 

opeved them before the fuel cells were starred of oxygen. This wonld have siiiiplified subsequent actions. 

38. Finding 

cate closed unless both the hydrogen and oxygen valves are closed. 

The fuel cell reactant valve talkback indicators in the spacecraft do not indiD etcriiziii ations 

(1) If these talkbacks vere designed so that either a hydrogen or oxygen valve 

closure would indicate “barberpole,” the Apollo 13 crew could po%il)ly hare 

acted in time to delay the failure of fuel cells 1 and 3. although they would nevertheless have failed when oxygen tank no. 1 ceased to supply oxygen. 

12) The ultimate outconie would not have been changed. but had the fuel cells 

not failed, Mipsion Control ancl the crew would not hare had to contend nith 

the failure of dc main bus B and ac bus 2 or attitude control problems while 

trying to evaluate the situation. 
REACTION COPI'TROL SYSTEM 

SD. Findings 

(a) The crew reported the talkback indicators for the heliuni isolation valres 

in the SJI RCS quads B aiid D indicated closed shortly after the dc main 

bus B failure. The seco'ndary fuel pressurization valves for quads A and C also 

were reported closed. 

(b) The SJI RCS quad D propellant tank pressures decreased until shortly 

after the crew was requested to confiriii that the heliuni isolation valves were 

aliened by the crew. 

(c) During the lx-hour period following the accident, JIission Control iiotecl 

that SJI RCS quad C propellant was not being used, although numerous firing 

signals were being sent to it. 

(d) Both the valve solenoids and the onboard indications of valve position 

of the propellant isolation valves for quad C are powered by dc main bus B. 

(e) During the 1YJ-hour period ininiediately following the accident, Mission 

Control advised the crew nhich SJI RCS thrusters to power and which ones to 

unpmver. 

Deterni.innt ions 

(1) The following valves were closed by shock at the time of the accident: 

Helium isolation ralres in quads B and D 

Secondary fuel pressurization valves in quads A and C 

(2) The propellant isolation valves in quad C probably were closed by the 

(3) Mission Control correctly deterniined the status of the RCS system and 

same shock. 

properly advised the crew on how to regain automatic attitude control. 

IIANAGEMENT OF ELECTRICAL SYSTEhI 

40. Findings 

(a) After fuel cell 1 failed. the total dc main bus -4 load was placed on fuel 

cell 2 arid the voltage dropped to approximately 2.5 volts, causing a caution and 

warning indication aiild a master alarni. 

(b) After determining the fuel cell 2 could not supply enough power to dc 

main bus A to maintain adequate uoltage, the crew connected entry battery A 

to this bus as an emergency measure to increase the bus voltage to its normal 

operating value. 

(c) Mission Control directed the crew to rediice the electrical load on dc main 

bus A by following the emergency powerdown clieclrlist contained in the onboard 

Flight Data File. 

(d) When the power requireiiients were sufficiently reduced so that the one remaining fuel cell would maintain adequate bus voltage, Mission Control directed the crew to take the entry battery off line. 

(e) Mission Control then clirected the crew to charge this battery in order to 

get as iiiiicli energy back into it as possible. before the ineyitable loss of the one fmictioning fuel cell. 

Deternr iiintions 

(1) Emergency we of the entry battery helped prevent Iiotential loss of dc 

iiiain bus A. wliicli could have led to loss of coniinunications between spacecraft 

and ground and other vital CM functions. 

(2) Available eniergency powerdown lists facilitated rapid reduction of loads on the fuel cell and batteries. 

-4TTEMPTS TO RESTORE OSYGES PRESSURE 

/tl. Findings 

iu) After determining that the CJI problems vere not due to instrumentation 

nialfnnctions. and after teiii1)orarily securing a stable electrical system coiifiguration. Jlission Control sought to improve oxygen 1)ressure by energizing the 

fan and lieatrr circuits in liotli oxygen tanks. 

(b) TT'lien tliese lirocedurrs failed to arrest the oxygen loss. Jlission Control 

directed the crew to shut down furl cells 1 and 3 by closing the hydrogen and 

oxygen flow valves. 
16 

Det erntinations 

(1) Under more normal conditions oxygen pressure might have been increased 

by turning on heaters and fans in the oxygen tanks ; no other known actions had 

such a possibility. 

(2) There was a possibility that oxygen was leaking downstream of the ralres ; 

had this been true, closing of the valves might have preserved the remaining 

oxygen in oxygen tank no. 1. 

42. Findings 

(a) Vith imminent loss of oxygen from oxygen tanks no. 1 and no. 2. and 

failing electrical power in the CAI. it was necessary to use the lunar module 

(LN) as a "lifeboat" for the return to Earth. 

(b) Mission Control and the crew delayed LJI actiration until about 15 

minutes before the SJI oxygen supply was depleted. 

(c) There were three different LN actiration checklists contained in the 

Flight Data File for normal and contingency situations : however, none of these 

was appropriate for the existing situation. It was necebsary to actirate the LJ1 as rapidly as possible to conserre LJI consumables and CAI reentry batteries to 

the maximum extent possible. 

(d) Mission Control modified the normal LJI actiration checklist and referred 

the crew to specific pages and instructions. This bypassed unecessary steps and 

reduced the activation time to less than an hour. 

(c) The LA1 inertial platform was aligned during an onboard checklist proce- dure which manually transferred the GI1 alignment to the L3L 

D eterni ina tions 

(1) Initiation of LRI activation was not undertaken sooner because the crew n7as properly more concerned with attempts to conserve remaining SJI oxygen. 

(2) Mission Control was able to make workable on-the-spot modifications to 

the checklists which sufficiently shortened the time normally required for powering up the Lhl. 

43. Findings 

(a) During the LM powerup and the CSM powerdown, there was a brief time 

interval during which Mission Control gave the crew directions which resulted 

in neither module having an active attitude control system. 

(b) This caused some concern in Mission Control because of the possibility of the spacecraft drifting into inertial platform gimbal lock condition. 

(c) The Command Module Pilot (CJIP) stated that he was not concerned 

because he could have quickly reestablished direct manna1 attitude control if it 

became necessary. 

LL-SAR MODULE ACTJXATIOS 

Deternaination 

This situation vas not hazardous to the crew because had gimbal lock 

actually occurred, sufficient time was arailable to reestablish an attitude 

reference. 

44. Findings 

(a) LII flight controllers were on duty in Mission Control at the time of the 

accident in support of the scheduled crew entry into the LJI. 

(b) If the accident had occurred at some other time during the translunar 

coast phase, LJI system specialists would not hare been on duty, and it n-oulcl 

have taken at least 30 minutes to get a fully manned team in Xission Control. 

,4lthough LBI flight controllers were not required until more than an hour 

after the accident, it was beneficial for them to be present as the problelll developed. 

&. Bindlngs 

(a) The L3I Tvas designed to support tvo men on a 2-day exmdition to the 

lunar surface. hlission Control made major revisions in the use rate of water, 

oxygen, and electrical power to sustain three men for the 4-day return trip to 

the Earth. 

(h) An emergency poTverdown checklist was arailable in the Flight Data File on board the LJI. Minor revisions were made to the list to reduce electrhl 

Determination 

Lhf COSSLTMABLES MANAGEMEST 
17 

energS requirements b about 20 percent of normal operational values with a corresponding reduction in usage of coolant loop water. 

( c) JIission Control cletermined that this maximum po1verdon.n could be delayed until after 80 hours ground elapsed time, allowing the Lhl primary guid- ance and navigation ,system to be liept pm-ered up for the second abort maneuver. (d) JIission Control dereloped contingency plans for further reduction of LM 

pojl-er for nse in case an LJI battery problem developed. Procedures fur use of 

CJI water in the LJI also were developed for use if needed. 

(0) Tonyard the end of the mission, sufficient consumable margins existed to 

a110~1- usage rates to be increased above earlier planned levels. This npas done. 

(f) TVhen the LJI \3*as jettisoned at 141 : 30 the approximate remaining margins were : 

Electrical power. 41/? hours. 

Water, 5y2 hours. 

Oxygen, 124 hours. 

Dcternzinat ions 

I 1) Earlier contingency ~Alaris tnid available checklists were adequate to ex- tend life support capability of the LJI ivell beyond its normal intended capability. 

(2) Mission Control maintained the flexibility of being able to further increase 

the LJI consumables margins. 

MODIFICATIOS OF LM CARBOS DIOXIDE REMOVAL SYSTEhI 

46 Findings 

(a) The lithium hydroxide ( LiOH) cartridges, which remove water and car- bon dioxide from the LJI cabin atmosphere, would have become ineffective due 

to saturation at about 100 hours. 

(0) Mission rules set maximum allowable carbon dioxide partial pressure at 

73mm Hg. LiOH cartridges are normally changed before oabin atmosphere car- bon dioxide partial preissure reaches this ralue. 

(c) Manned Spacecraft Center engineers derised and checked out a procedure 

for using the Ci\I LiOH cannisters to achieve carbon dioxide removal. Instructions were given on hon- to build a modified cartridge container using materials in 

the spacecraft. 

(A) The crew made the modification at 93 hours, and carbon dioxide partial 

pressure in the LhI dropped rapidly from 73mm Hg to O.lmm Hg. 

(0). Jlission Control gave the crew further instructions for attaching additional 

cartridges in series with the first modification. After this addition, the carbon 

dioxide partial pressure remained below 2mm Hg for the remainder of the 

Earth-return trip. 

Determination 

The Manned Spacecraft Center succeeded in improvising and checking out a modification to the filter system which maintained carhon dioxide concentration 

well within safe tolerances. 

47. Findings 

(a) During the time interval between 97:13:53 and 97:13:66, LRI descent 

battery current measurements on telemetry showed a rapid increase from values 

of no more than 3 amperes per battery to ralues in excess of 30 amperes per battery. The exad ralue in one battery cannot be determined because the nieasurement for battery 2 was off-scale high at 60 amperes, (b)dt about that time the Lunar JIodule Pilot (LJLP) heard a “thump” from 

the vicinity of the LM descent stage. 

(c) When the LhlP looked out the La1 right-hand window, lie observed a venting of small particles from the general area where the L?II descent batteries 

1 and 2 are located. This venting continned for R few minutes. 

(d) Prior to 97:13 the battery load-sharing among the four batteries had been 

equal, but immediately after the battery currents returned to nominal, batteries 

1 and 2 supplied 9 of the 11 amperes total. By 97 :23 the load-cliaring had returned 

to equal. 

(6‘) There was no electrical interface between the Lhl and the CSM at this 

time. 

(f) An hISC investigation of the anomaly is in progress. 

LM ANOhCALY 
18 

Determination 

(1) An anomalous incident occurred in the LM electrical system at about 

(2) The thump and the venting were related to this anomaly. (3) The apparent short circuit cleared itself. 

(4) This anomaly was not directly related to the CSJI or to the accident. 

(5) This anomaly represents a potentially serious electrical problem. 

97 :13 553 which appeared to be a short circuit. 

031 BATTERY RECHARGING 

48. Findings 

(a) About one half of the electrical capacity of reentry battery A (20 of 40 

amp-hours) was used during emergency conditions following the accident. d small 

part of the capacity of the reentry batter B was used in checking out dc main bus 

B at 95 hours. The reduced charge remaining in the batteries limited the amount 

of time the CAI could operate after separation from the LJI 

(b) Extrapolation of LM electrical power use rates indicated a capacity in excess of that required for LM operation for the remainder of the flight. 

(c) Mission Control worked out a procedure for using LM battery power to 

recharge CJI batteries A and B. This procedure used the electrical umbilical between the LN and the C31 which normally carried electrical energy from the CAI 

to the LX The procedure was nonstandard and was not included in checklists. 

(d) The procedure was initiated at 112 hours and CAI batteries A and B were fully recharged by 128 hours. 

Det emnin ation 

Although there is always some risk involved in using new, untested procedures. 

analysis in advance of use indicated no hazards were involved. The procedure 

worked very well to proride an extra margin of safety for the reentry operation. 

49. Findings 

(a) After the accident, it became apparent that the lunar landing could not be 

accomplished and that the spacecraft trajectory must be altered for a return to 

Earth. 

(a) At the time of the accident, the spacecraft trajectory was one whieh would 

have returned it to the vicinity of the Earth, but it vould have been left in orbit 

about the Earth rather than reentering for a safe splashdown. 

(c) To return the spacecraft to Earth, the following midmurse corrections were made: 

A 38-fps oorrection at 61:30, using the LhI descent propulsion system (DPS), required to return the spacecraft to the Earth. 

An 81-fps burn at 79:28, after swinging past the Moon, using the DPS 

engine, to shift the landing point from the Indian Ocean to the Pacific and 

to shorten the return trip by 9 hours. 

h 7.8-fps burn at 105 :18 using the DPS engine to lower Earth perigee from 

87 miles to 21 miles. 

h 3.2-fps correction at 137 :40 using LM RCS thrusters, to assure that the 

CM would reenter the Earth’s atmosphere at the cenkr of its corridor. 

(d) A11 course corrections were executed vith expected accuracy aind the C31 

reentered the Earth’s atmosphere at 142 :40 to return the crew safely at 142 :54. 

near the prime recovery ship. 

(e) Without the CAI guidance and navigation system, the crew could not 

navigate or compute return-to-Earth maneuver target parameters. 

TBAJECTORY CHANGES FOB SAFE RETURN TO EARTH 

Determinations 

(1) This series of course corrections was logical and had the best chance of success because, as compared to other options, it avoided use of the damaged SJI : 

it put the spacecraft on a trajectory, within a fern hours after the accident, which 

had the best chance for a safe return to Earth : it placed splashdon-n where the 

best recovery forces were located ; it shortened the flight time to increase safety margins in the use of electrical power and water; it conserved fuel for other course corrections which might have become necessary ; and it kept ope11 an optioll 

to further reduce the flight time. 

(2) Mission Ctontrol trajectory planning and maneuver targeting were essen- tial for the safe return of the crew. 
19 

ENTRY PROCEDURES AND CHECKLISTS 

50. Pindings 

(a ) Preparation for reentry required nonst iiidard procedures because of tlie 

lack of SAI oxygen and electrical power supplies. 

(b) The SJI RCS engines iioriiiallr providr separation betiveen the SRI and 

the CJI by continuing to fire after separation. 

(c) Apollo 13 SJl RRCS engiiies could not contiiiue to flre after separation 

because of the earlier failure of the fuel cells. 

(d) Tlie CAI guidance and navigation system was powered down due to the 

accident. The LA1 guidance and navigation sptem had also been powered down 

to conserve electrical energy and water. A spacecraft inertial attitude reference 

had to be established prior to reentry. 

(c) The reentry preparation time had to be extended in order to accomplish 

the additional steps required by the unusual situation. 

(f) In order to conserve the CJI batteries, LBI jettison mas delayed as long as practical. The LRI batteries mere used to supply part of the power necessary 

for CAI activation. 

(3) Tlie procedures for accoiiiplishiiig the final course correction and the re- entry preparation were developed by operatiims support personnel under the 

direction of Jlissioii Control. 

(h) An initial set of procedures was defined within 12 hours after the accident. These were refined and modified during the following 2 days, and evaluated 

in siiiiulators at NSC and KSC by members of the backup crew. 

(i) The procedures Tvere read to the crew about 24 hours prior to reentry, allowing the crew time to study and rehearse thm. 

(j) Trajectory evaluations of contingency cinditions for LA1 and SJI separa- tion were conducted and documented prior to the niission by mission-planning personnel at JISC. 

(1;) Most of the steps taken were extracted from other procedures which had 

heen developed. tested, and simulated earlier. 

Determinations 

(1) The procedures developed worked well and generated no new hazards 

beyond those unavoidably inherent in using procedures which have not been care- fully developed, simulated, and practiced over ,I long training period. 

(2) It is not practical to develop, simulate, and practice procedures for use in every possible contingency. 

5i. Findings 

(a) During the reentry preparations, after SBI jettison, there was a halfhour period of very poor coniinunications wilh the CBI due to the spacecraft 

being in a poor attitude\vitli the LN present. 

(6) This condition mas not recognized by the crew or by Jlission Control. 

Determination 

Soiiie of tlie reentry preparations were unnecessarily yroloiiged by the poor 

communications. but since the reentrx preparztion time-line was not cro\viled. 

the delay was more of a nuisance than an additional hazard to the crew. 

52. Findings 

(a) The crew maneuvered the spacecraft to the wrong LA1 roll attitude in 

preparation for LJI jettison. This attitude put tlie CJI very close to ginilial lock 

which. liad it occurred. ~~ould have lost the inertial attitude reference essential 

for an automatic guidance system control of reeiitry. 

(b) If gimbal lock liad occurred, a less accurate but adequate attitude refer- ence could have been reestablished prior to reent ry. 

Determination 

The nost significant consequence of losing the attitude reference iii this situation n-ould have been tlie subsequent impact nil the reniaininq reentry prepnrntiori timeline. In taking tlie time to reestalilisii this reference, less tiiiie lvonid 

have been available to acconilAisli the rest of the necessary 1JroCedUre5. Tlie occ~irPence of gimbal lock in itself would not have significantly increased the cren- hazard. 
PART 4. RECOMMEKDATIONS 

The cryogenic oxygen storage system in tlie service module should be modified 

a. Remove from contact with the oxygen all wiring. and the unsealed 

motors, which can potentially short circuit and ignite adjacent materials ; 

or otherwise insure against a catastrophic electrically induced fire in tlie 

tank. 

b. Minimize the use of Teflon. aluniinnni. and other relatively combustible 

materials in the presence of the oxygen and potential ignition sources. 

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

3. The warning systerns on board the dpollo spacecraft and in tlie Uission Colitrol Center should be carefully rerielved and modified where appropriate. \yith 

specific attention to the following : 

a. Increasing the differential between master alarm trip levels and 

expected normal operating ranges to avoid unnecessary alarms. 

b. Changing the caution and warning system logic to prevent an out-oflimits alarm from blocking another alarin when a second quantity in the same subsystem goes out of limits. c. Establishing a second level of limit sensing in Mission Control on critical quantities ivith a visual or audible alarnl which cannot be easily 

overlooked. 

d. Providing independent talkbnck indicators for each of the six fuel cell 

reactant \-alves plus a master alarni when any Tralve closes. 

4. Consuiiiables and emergency equipment in the LA1 and the CU should be 

reviewed to determine whether steps should be taken to enhance their potential 

for use in a “lifeboat” mode. 

6. The Manned Spacecraft Center should compete the special tests and analyses now underway in order to understand more coinpletely tlie details of the dnollo 

13 accident. In addition. tlie lunar niodule power systeni anonialies should receive 

careful attention. Other SASA Centers should continue their support to MSC in 

the areas of analysis and test. 

6. Whenever significant anoiiialies occur 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 

hare previously been corrected or explained. Furthermore, critical decisions involving the flightworthiness of subsystems should require the presence and full 

participation of an expent who is intimately familiar with the details of that 

subsystem. 7. NASA should conduct a thorough reexamination of all of its spacecraft. 

launch vehicle, and ground systems which contain high-density oxygen, or other 

strong oxidizers, to identify and evaluate potential combustion hazards in the 

light of inforniation developed in this investigation. 

8. SASd should conduct additional research on materials compatibility, ignition, and combustion in strong oxidizers at various g levels; and on tlie characteristics of supercritical fluids. Where appropriate, new SBSh design standards 

should be developed. 9. The Manned Spacecraft Center should rea all Bpollo spacecraft subsyhtenis, and the engineering organizations respo for them at JISC and at itc 

prime contractors, to insure adequate understanding and control of the engineering and manufacturing details of these subsystems at the subcontractor and 

yelldor level. \vhere necessary, organizational elenients should be strengthened 

and in-depth reyie\vs conducted of selected aubsysteills with eniphasis 011 Soundness of design, quality of manufacturing, adequacy of test, and operational 

experience. 

to : 
21 

STATEMENTS OF DiR. THOMAS 0. PAINE, ADMINISTRATOR OF NASA: 

EDGAR M. CORTRIGHT, DIRECTOR, LANGLEY RESEARCH CENTER 

AND CHAIRMAN OF THE APOLLO 13 REVIEW BOARD; DR. 

CHARLES D. HARRINGTON, CHAIRMAN, AEROSPACE SAFETY 

ADVISORY PANEL ; DR. DALE D. MYERS, ASSOCIATE ADMINISTRATOR FOR MANNED SPACE PLIGHT; AND DR. ROCCO A. PETRONE, 

APPOLLO PROGRAM DIRECTOR 

Dr. PAINE. I would like to ask Mr. Cortright to begin the testilnony 

this morning, Mr. Chairman, by giring a brief summary of the report 

of the hpollo 13 Review Board. 

(Biographical data of the witnesses appear at the end of this 

hearing.) 

STATEMEST 131- JIR. CORTRIGHT 

Mr. CORTRIGHT. Bfr. Chairman, ineinbers of the committee, with your 

permission I would like to submit for the record today a statement 

and summarize it for the committee. The statement recounts in some 

detail the establishment aiid operation of the Apollo 13 Review Board 

including the extensive test program conducted for the Board. 

SUJIMARP OF BOARD'S REPORT 

The Board's report vhich TT~S submitted to the Administrator on 

,June 15 aiid copies of which were submitted to this coinmittee on the 

same date contains over 30 pages of finclings and cleterminations. It is 

these findings and determinations which I would like to summarize 

for you this morning by reading from the introduction to chapter 5 of 

the report. 
To 

cel 

1 Overboard 

FIGURE 1. Oxygen tank No. 2 internal components. 

Slide 1 (see fig. 1) shows the simplified drawing of the oxygen 

storage tank in which the accident occurred. I will just say a brief 

word about this slide to orient you to the nature of the problem that 

was encountered. 

The tank itself was a high-pressure vessel which contains oxygen 

in a supercritical high density state. The tank is made of high strength 

Inconel. It is a doubled wall tank. The inner wall carries the pressure 

and the outer wall is there for insulation purposes. There is insulation 

between the two walls. 

Sow, the two major assemblies within the tank are a quantity gage 

which is shown on the left and a heater-fan assembly rd-hich is shown 

on the right. The problem as I will describe shortly, occurred primarilr 

with the heater-fan assembly which was overheated and damaged. 

I will come to the next slide in a moment. 
23 

On the table in front of me are the quantity probe itself and the 

heater-fan assembly which I will be happy to show you after the 

hearing if you care to look at it. 

In addition, on my left is a cut open tank which i~as subjected to a 

fire simulating that which actually occurred during the mission, and 

this tank is available for your examination. 

Nm-, in brief, this is whnt happened. After assembly and acceptance 

testing the oxygen tank No. 2 which flew on A4po110 13 was shipped 

from Beech Aircraft Corp. to North American Rockwell in apparently 

satisfactory condition. It is now known, however, that the tank contained two protective thermostatic switches on the heater assembly. 

These switches were inade uate and subsequently failed during ground 

test operation at Kennedy 1 pace Center. 

Teliiperatiire seiisor 

Quailtity probe - 

FIGURE 2.Oxygen tank wiring and lines. 
24 

Now, the second slide (see figure 2) shows this in a little more 

detail. As you can see at the top of the heater-fan assembly the word 

“thermostat” is shown with two arrows. These arrows point to the 

two thermostatic siTitches which failed. In addition, there is a fill tube 

assembly shown right here at the top of the quantity prdbe which was 

loose in a niaiiner which I will describe in just a moment and this led 

to a special detailking procedure which failed those switches and ultimately damaged the wiring. 

Now, in addition to these thermostatic switches which subsequently 

failed, it is probable that the tank contained a loosely fitting fill tube 

assembly which I just pointed out. This assembly mas probably displaced during subsequent handling after shipment and this handling 

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 the use of improvised detailking procedures at the 

Kennedy Space Center nThich almost certainly set the stage for the 

accident. 

NOT, although Beech did not encounter any problem in detailking 

during the acceptance test of this tank, it was not possible to detank 

the oxyeii tank No. 2 using normal procedures at the Kennedy Space 

Center. Tests and analyses indicate that this was due to gas leakage 

through this displaced fill tube assembly that I mentioned. 

Now, the special detailking procedures at Kennedy 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 tests for the conditions experienced. However, the procedures did not violate the specifications which govern the operation 

of the heaters at the Kennedy Space Center. 

In reviewing these procedures before the flight. officials of N-4SB. 

North American Rockwell, and Beech did not recognize hhe possibility 

of damage due to overheating. Many of these officials were not aware 

of the extended heater operation. In any event, the thermostatlc 

switches might have been expected to protect the tank. 

A number of factors contribiited to the presence of inadequate 

thermostatic switches in the heater assembly. The original 1962 

specifications from North American Rockwell to Beech Aircraft Corp. 

for the tank and heater assembly specified the use of 28 volt D.C. po1Ter 

which is used in the spacecraft. In 1965 Korth American Rockwell 

issued a revised specification which stated that the heater shonld use 

a 65 volt d.c. power supply for tank pressurizaiion. This was the poJT7er 

supply used at Kennedy to reduce pressnrization time. Beech ordered 

switches for the block 2 tanks but did not change the sdch s13ecifications to be compatible with 65 volt D.C. 

The thermostatic switch discrepancy was not detected by NA1SA4, 

North American Rockwell or Beech in their review of documentation, 

nor did tests identify the incompatibility of the sm4tches with the 

promid support equipment at Kennedy. since neither qualification nor 

acceptance testing required switch cycling under load as should hare 

been done. It vas a serious oversight in n-hich all parties shared. 

Thermostatic switches could accommodate the 65 rolt D.C. during 

tank pressurization, however, because they normally remain cool and 

closed. HoTvever, they conld not open without damage with the 65 
25 

volt d.c. power supply. They mere never required to do SO until the 

special detanking. 

During this procedure as the switches started to open when they 

reached their upper temperature-they were welded permanently 

closed or otherwise failed pernlalielltly closed by the resulting arc and 

JTere rendered inoperative as protective thermostats. 

Kow, the next slide (see fig. 3) shows a photograph of a switch 

FIGURE 3 

which was failed vith this current during tests at the Manned Spacewaft Center. The drawing is in the report and in the statement that 

I submitted today-the photograph is in there. 

Failure of the thermostatic switches to open could hare been detected 

at Kennedy Space Center if switch operation had been checked by 

observing heater current readings 011 the oxygen tank heater control 

panel. Although it was not recognized at that time, the tank temperntnre readings indicated that the heaters had reached their temperature 

limit and svitch openings shonld hare been expected. As shovn by 

subsequent tests, failure of the thermostatic switches probably permitted the temperature of the heater tube assembly to reach about a 

1,000 degrees Fahrenheit in spots during the continuous eight-hour 

period of heater operation. Such heating has been shown by tests 

to damage severely the Teflon insulation on the fan motor wires in 

thc vicinity of the heater assembly as shown in the next slide. This is 
26 

FIGURE 4 

a picture of wiring which was taken from the heater assembly after 

the simulated tank operation in a nitrogen environment. Had this 

been in an oxygen environment other tests hare shown that the 

insulation deterioration can be even worse. (See fig. 4.) 

From that time on, including pad occupancy, the oxygen tank No. 2 

was in a hazardous condition when filled with oxygen and electrically 

powered. It T~S not until nearly 56 hours into the mission, howerer, 

that the fan motor wiring, possibly moved by the fan stirring of the 

contents of the tank, shortcircuited and ignited this insulation by 

means of an electric arc. The resulting combustion in the oxygen tank 

probably overheated and failed the wiring conduit where it enters the 

tank. This is the tube up here [iiidicating] where it goes into the top of 

the tank, and possibly a portion of tlie tank itself, primarily the cap 

that goes through the tank at this point. 

The rapid expulsion of high pressure oxygen which followed, possibly augmented by combustion of the insulation in the space surrounding the tank, blew off the outer panel to bay KO. 4 in the service module, 

caused a leak in the high pressure system of oxygen tank No. 1, damaged the high gain antenna, caused other iniscellaneous damage and 

aborted the mission. 

The accident is judged to have been nearly catastrophic. Only outstanding performance on the part of the crevi-, mission control and 

other members of tlie team which supported the operation successfully 

returned the crew to earth. 

Now. in investigating the accident to Apollo 13, Bfr. Chairman, the 

Board has also attempted to identify those additional technical and 
27 

lnanagement lessons which can be applied to help assure the success 

of future space flight missions. Several recommendations of this nature 

are included. In addition, I would like to say that the Board recognizes 

that the contents of this report are largely of a critical nature. The 

report highlights in detail faults or deficiencies in equipment and procedures that the Board has identified. This is the nature of a review 

board report. 

It is important, liowerer, in our judgment, to deal with criticisms of 

this report in a broader context. The Apollo spacecraft system is not 

without shortcomings but it is the only system of its type ever built 

and successfully demonstrated. It has flown to the moon five times and 

landed t\.i.ice. The tank which failed, the design of which is criticized 

in this report, is one of 51 series which has thousands of llours of successful operation in space prior to Apollo 13. 

In addition, while the team of designers, engineers and technicians who have built and operate the Apollo spacecraft also has its 

shortcomings, the wcomplishinents speak for themselves. We feel by 

hardheaded criticism and continued dedication this team can maintain this Nation’s preeminence in space. 

Thank you very much. 

(Mr. Cortright’s prepared statement follo1.i.s :) 

STATEMENT OF EDGAR 31. CORTRIGHT, CHAIRMAN, APOLLO 13 REVIEW BOARD, 

NATIONAL AERONACTICS AKD SPACE ADJIINISTRATION BEFORE THE COX MIWEE ON 

~~ROXAUTICAL AKD SPACE SCIEXCES, U.S. SENATE 

Yr. Chairman and Members of the Committee: I appreciate this opportunity 

to appear before the Committee to’sunimarize the Report of the Apullo 13 Review 

Board. 

As you know, I presented this Report on behalf of the Board to the Administrator and Deputy Administrator on June 1.5. 1970. At that time, copies of the 

Report were given to the Members and Staff of the Committee, and the Report was made public. This morning I would like first to outline for the Committee how the Board was established and hom- it organized itself to review and report on the Apollo 13 

accident. Then I will cover in some detail the findings and determinations of the 

Board regarding the accident, including pre-accident mission ei7ents, the eyents 

of the accident itself, and the recovery procedures which were iniplemented to 

return the crew safely to earth. I rill also summarize the Board’s findings and 

determinations regarding the management, design, manufacturing, and test procedures employed in the Apollo Program as they relate specifically to the accident. 

Based on its findings and determinations, the Board made a series of detailed 

recommendations. These are set forth at the end of my statement. 

ESTABLISHIIENT AND HISTORY OF THE BOARD 

The Apollo 13 Review Board was established, and I was appointed Chairman, on April li, 1970. The charter of the Board mas set forth in the memorandum 

which established it. ‘Cnder this charter the Board was directed to : 

“(a) Review the circumstances surrounding the accident to the spacecraft ivhich occurred during the fiight of Apollo 13 and the subsequent fiight and 

ground actions taken to recover, in order to establish the probable cause or causes of the accident and assess the effectireness of the recovery actions. 

‘‘ (b) Reriew all factors relating to the accident and recovery actions the Board 

determines to be significant and relevant, including studies, findings, recom- niendations. and other actions that hare been or may be undertaken by the program offices. field centers. and contractors involved. “ (c) Erect such further specific investigations as may be necessary. 

“(d) Report as soon as possible its findings relating to the cause or causeq of the accident and the effectireness of the flight and ground recovery action&. 

“ (e) Develop recomniendations for corrective or other actions, based upon its findings and determinations or conclusions clerivd therefrom, “ (f) Docuinent its findings, determinations. and recommendations and submit a final report.” 
The Umbership of the Board n'as established on april 21, 1970. The memhIr. Edgar JI. Cortright, Chairman (Director, Langley Research Center) 

Mr. Robert F. Bllnutt (Assistant to the Administrator, SASA Hqs) 

Jlr. Seil Armstrong (Astronaut, Manned Spacecraft Center) 

Dr. John F. Clarli (Director, Goddarcl Space Flight Center) 

Brig. General Walter R. Heclriclr, Jr. (Director of Space, DCS/RkD. Hgs 

hIr. Vincent L. Johnson (Deputy Associate Administrator, Engineering, OfNr. llilton Klein (Jlaiiager, -XEC-ShSA Space Suclear Propulsion Office) 

Dr. Hans AI. Mark (Director, hies Research Center) 

Legal Counsel to the Board is AIr. George T. Jlalles, Chief Counsel, Laiigley 

Appointed as Obqervers were : 

JIr. TS'illiam A. Anders i Ewxutire Secretary, Satioiial heronantics and 

Dr. Charles D. Harrington (Chairman. SASA Aerospace Safety Advisory 

Mr. I. I. Pinkel (Director, Aerospace Safety Research and Data Institute, 

11r. James E. Wilson, Jr., (Technical Consultant, House of RepresentaThe documents establishing the Board and its membership and other relerant documents are included in Chapter 1 of the Board's Report. 

The Review Board convened at the Manned Spacecraft Center (JISC), Honston, Texas, on Tuesday, April 21, 1970. Four Panels of the Board were formed, 

each under the orerview of a member of the Board. Each of the Panels nns 

chaired by a senior official experienced in the area of review assigned to the 

Panel. In addition, each Panel was manned by a number of experienced specialists to provide in-depth technical competence for the review actiI*itF. During 

the period of the Board's activities, the Chairmen of the four Panels \?*ere 

responsible for the conduct of reviews, evaluations, analyses, and other studies 

bearing on their Panel assignments and for preparing documented reports for 

the Board's consideration. Complementing the Panel efforts, each member of 

the Board assumed specific responsibilities related to the overall review. 

bers are: 

USAF) 

fice of Space Science and dpplications) 

Research Center. 

Space Council) 

Panel) 

KA4SA Lewis Research Center) 

tives, Committee on Science and Astronautics) 

APOLLO 13 I REVIEW BOARD 

E M CORIRIGHT. CHAIRMAN 
On figure 6 is qhown a chart depicting the organization of the Board. The 

four Panels-Mission Events, Manufacturing and Test, Design, and Project 

Management-are shown along with the subpanels and the supporting office 

structure. The membership and responsibilities of each Panel are set forth in 

the Report. 

While the Board’s intensive review actirities were underway, the Manned 

Spacecraft Center dpollo 13 Investigation Team, under James A. McDivitt, Director of the JISC dpollo Spacecraft Program Office, was also conducting its own analysis of the Apollo 13 accident. Coordination betn-een the Investigation 

Team work and the dpollo 13 Rerieiv Board actirities ivas effected through the 

JIannecl Space Flight Technical Support official and by maintaining a close and 

continuing working relationship between the Panel Chairmen and officials of 

the JISC Investigation Team. In addition. Board inenibers regularly attended 

daily status meetings of the JIanned Spacecraft Center Investigation Team. 

In general, the Board relied on Manned Spacecraft Center post-mission evaluation activities to provide the factual data base for evaluation, assessment. and 

analysiq efforts. However, the Board, through a regular procedure, also levied 

specific data collection, reduction, and analysis requirements on MSC. Test 

support for the Board was provided by JISC, but in addition, the Board established an extensive series of special tests and analyses at other NASA Centers 

and at contractor facilities. JIeinbers of the Board and its Panels also visited 

contractor facilities to review manufacturing, assembly, and test procedures applicable to Apollo 13 mission equipment. 

In this test program, which included nearly 100 separate tests. and whioli involved several hundred people at its peak, the elenients of the inflight accident were reproduced. All indications are that electrically initiated combustion of‘ 

Teflon insulation in oxygen tank KO. 2 in the service module was the cause o€ the Apollo 13 accident. One series of tests demonstrated electrical ignition 

of Teflon insulation in supercritical oxygen under zero g and at one g, and provided data on ignition energies and burning rates. Other tests, culminating in a complete flight tank coiiibustion test. denionstrated the most probable tank 

failure mode. Simulated tank rupture tests in a W scale service module verified 

the pressure levels necessary to eject the panel from the service module. Other 

special tests and analyses clarified how they might hare been generated. I have 

with me a brief film, highlighting these tests, which I would like to shonr at 

the conclusion of my statement. 

MOLL0 13 SYSTEM6 

Before tracing the analyses which led to the Board’s conclusioas-and to place them in proper context-I mould like to explain the design and functions of the 

oxygen tank #2 as a part of the Apollo system. Details of the entire Apollo/ 

Saturn Space Vehicle are set forth in the Report and its Appendices. 

47476 0-7- 
Service !iioaule 

(pacecraft LM adapter 

Spacecraft 

30 

Space veliicle 

FIGURE 6 

ft 

LAUNCH VtHlCLl 

Laiiiicli vetiicle 

Figure 6 shows the hpollo/Saturn Space Vehicle, with which you are all 

familiar. Figure 7 shows the service inodule which, as you know, is designed to 

provide the main spacecraft propulsioii and maneuvering capability during a 
31 

SECTOR 2 

SECTOR 3 

SECTOR 4 

SECTOR 5 

SECTOR 6 

CENTER SECTION SERVICE PROPULSION EHGINE AN0 

HELIUM TANKS 

FIGURE 7 

mission. It also contains most of the spacecraft consuinables (oxygen, water, 

propellant, and hydrogen) and supplies electrical power. The service module is 

divided into six sectors or hays surrounding a center section. Tlie oxygen tank, 

to which I referred, is located in Bay 4 (shown in more detail on figure 8), 
32 

NASA-S -70-512-1' 

ARRANGEMENT OF FUEL CELLS AND CRYOGENIC 

C,-d/ 

FIGURE 8 

along with another oxygen tank, tmo hydrogen tanks, three fuel cells and interconnecting lines, and ineasuring and control equipment. The tanks supply oxygen 

to the enriromnental control systeni (ECS) for the astronauts to breathe, and 

oxygen and hydrogen to the fuel cells, The fuel cella geiierate the electrical 

power for the conmiand and service iiiodules during a mission. The nest slides 

(figures 9, 10, and 11) are photographs of Bay 4 of the serrice module for Apollo 

13, showing the major elements and their interconnection. Slide 7 shows the 

oxygen tank #2 in place. 
33 

FIGURE 9 
34 

FIGURE 10 
FIGURE 11 

As the simplified drawing in the next slide indicates (see figure 12) each 

oxygen tank has an outer shell and an inner shell, arranged to proride a vacuum space to reduce heat leak, aiid a dome enclosing paths into the tank for transmission of fluids, and electrical power and signals. The space between the shells 

and the space in the dome are filled with insulating materials. Mounted in the 

tank are two tubular assemblies. One, called the heater tube, contains two themlostatically protected heater coils and two sniall fans driven by 1800 RPJL motors 

to stir the tank contents. The other assembly, called the quantity probe, consists 

of a cylindrical capacitance gage used to measure electrically the quantity of 

fluid in the tank. The inner cylinder of this probe is connected through the top 
36 

To fuel 

C 

Pressure 

transduce 

L valve I ITverboard 

J?IQURE 12 

Closeout cap 

of the tank to a fill line from the exterior of the SM and serves both 'as a fill and 

drain tube and as one plate of the capacitance gage. In addition, a temFrature sensor is mounted on the outside of the quantity probe near the head. Wiring 

for the quantity gage, the temperature sensor, the fan motors, and the heaters 

passes through the head of the quantity probe, through a conduit in the dome 

and to a connector to the appropriate external circuits in the CSJI. The routing 

of wires and lines from the tank through the dome is shown in slide 9 (see fig- ure 13). 

The oxygen tank, as designed, contained materials, which if ignited will burn 

in supercritical oxygen. These include Teflon, used, for examiple, to insulate the 

wiring, and aluminum. 

Pressure in the tank is measured by a pressure gage in the supply line, and a pressure smitch near this gage is provided to turn on the heaters in the oxygen 

tank if the pressure drops below a preselected value. This periodic addition of heat 

to the tank maintains the pressure at a sufficient level to satisfy the demand for 

oxygen as tank quantity decreases during a flight mission. 
37 

Teniperature setisor 

Quaiitlty probe - 

iostat 

FIGURE 13 

The oxygen tank is designed for a capacity of 320 pounds of supercritical 

oxygen at pressure3 ranging between 865 and 936 pounds per square inch absolute 

(psia). The tank is initially filled with liquid oxygen at -297°F and operates 

over the range from -340°F to +SOOF. The tenii "supercritical" means that the 

oxygen is maintained at a temperature and pressure which assures that it is a homogenous. single-phase fluid. 

The burst precsure of the oxygen tank is ahut 2200 psia at -1.5O'F. over twice the nornial operating grrbsure at that teiniierature. A relief valve in the 

supply line leading to the fuel cells and the ECS is designed to relieve pressure 

in the oxygen tank at a presaure of approximately 1000 psi. The oxygen tank 

dome is open to the vacuum between the inner and outer tank shell and contains a rupture disc designed to blow out at about 75 psi. 
38 

AS shown in figure 13, each heater coil is protected with a thermostatic switch, 

mounted on the heater tube, which is intended to open the heater circuit when it senses a temperature of 80°F. As I mill point out later in tracing the Board’s con- clusions as to the cause of the accident, when the heaters were powered from a 65 volt DC supply at KSC during an improvised detanking procedure, these 

thermostatic switches, because they were rated at only 30 V DC, could not prevent 

an overheating condition of the heaters and the associated wiring. Tests con- ducted for the Board indicate that the heater tube assembly was probably heated 

to a temperature of as much as 1000°F during this cletanking procedure. 

‘THE APOLLO 13 JfISSIOR* 

With this general background, I will noir- suininarize the Apollo 13 mimion. 

This mission, as you know, was designed to perfonn the third manned lunar 

landing. The selected site was in the hilly uplands of the Fra JI’auro formation. 

A package of five scientific experiments was planned for emplacement on the 

lunar surface near the lunar niodule landing point. Additionally the ilpollo 13 

landing crew was to gather the third set of selenological samples Gf the lunar 

surface for return to earth for extensive scientific analysis. Candidate future 

landing sites were wheduled to be photographed froin lunar orbit. The crew con- sisted cf Captain Janie A. Lovell, Commander; Fred W. Haise, Lunar Jlodule 

Pilot ; and John L. Smigert, Jr., Command Jlodule Pilot, who replaced Thomas 

I<. JIattingly, 111, who had been exposed to rubella and. after tests, found not to 

be immune. 

Launch was on time at 2:13 p.in., EST an April 11 from the KSC Launch 

Complex 39A. The spacecraft was inserted intlo a 100-nautical mile circular 

earth orbit. The only significant launch phase anomaly wm preiiiature shutdown 

of the center engine of the S-I1 second stage. This anomaly, although serious, 

was not related to the subsequent accident. It is being investigated by the Apollo 

organization. As a result of this shutdown, the remaining four S-I1 engines 

burned 34 seconds longer than planned and the S-IVB third stage engine burned a few seconds longer than planned. At orbital insedion, the velocity was within 

1.2 feet per second of the planned velocity. Moreover, an adequate propellant 

margin mas maintained in the S-IVB for the translunar injection burn. 

After spacecraft systems checlroat in earth orbit, the S-IVB restarted for 

the translunar injection (TLI) burn, with shutdown coming some six minutes 

later. After TLI, Apollo 13 was on the planned free-return trajectory with a predicted closest approach to the lunar surface of 210 nautical miles. 

The command and service module (CSJI) was separated from the S-IVB 

about three hours into the mission, and after a brief period of station-keeping, 

the crew maneuvered the CSM into dock with the LJI vehicle in the LJI adapter 

atop the S-IVB stage. The S-IVB stage x%s separated from the docked CSJI 

and LJI shortly after four hours into the mission, and placed on a trajectory 

to ultimately iiiipact the moon near the site of the seismometer emplaced by 

the A~ollo 12 crew. At 30 :40 :49 g.e.t. (ground elapsed time) a inidcourse correction maneuver was made using the service module propulsion system. This maneuver took 

Apollo 13 off a free-return trajectory and placed it on a non-free return trajectors. A similar profile had been flown on Apollo 12. The objective of leaving a free-return trajectory is to control the arrival time at the moon to insure the 

proper lighting conditions at the landing site. The transfer inaneuver lowered 

the predicted closest approach to the moon, or pericynthion altitude, from 210 to 

64 nautical miles. 

From launch through the first 46 hours of the mission. the performance of the 

oxygen tank #2 was normal, so far as telemetered data and crew observations 

indicate. At 46 :40 :02, the crew turned on the fans in oxygen tank #2 as a routine 

operation, and the oxygen tank #2 quantity indication changed from a normal 

reading to an obviously incorrect reading “off scale high” of over 100 percent. 

Subsequent events indicate that the cause was a short circuit which was not 

hazardous in this case. At 47 :34 :50 and at 51 :07 :44 the oxygen tank #2 fans were turned on again. 

with no apparent adverse effects. The quantity gage continued to read “off scale 

high.” 

Following a rest period. the Apollo 13 crew began preparations for activating 

and powering up the lunar module for checkout. At about 53 and one-half hours 
39 

g.e.t. Astronauts Love11 and Haise were cleared to enter the LA1 to commence in-flight inspection for the LJI. After this inspection period, the lunar module was powered down and preparations were underway to close the LhI hatch and run through the presleep checklist whtw the accident in oxygen tank #2 

occurred. 

At about 5.5:.33. flight controllers in the Jlission Control Center at JISC requested the crew to turn on the cryogenic system fans and heaters, since a master alarm on the CAI Caution and Warning System had indicated a low 

pressure condition in the cryogenic hydrogen tank #l. This tank had reached the 

low end of its normal operating pressure range several times previously during 

the flight. Swigert acknowledged the fan cycle request and data indicate that 

current was applied to the oxygen tank #2 fan motors at 55 :53 :20. 

About 2v2 minutes later, at 55 54 53.5, telemetry from the spacecraft was lost 

almost totally for 1.8 seconds. During the period of data loss, the Cantion and 

Warning System alerted the crew to a low voltage condition on DC Main Bus B, one of the two main buses which supply electrical power for the command module. 

At about the same time, the crew heard a loud “bang” and realized that a problem existed in the spacecraft. It is now clear that oxygen tank #2 or its 

associated tubing lost pressure integrity because of combustion within the tank, 

and that the effects of oxygen escaping from the tank caused the removal of the 

panel covering Bay 4 and a relatively slow leak in oxygen tank #1 or its lines or valves. Photographs of the service module taken by the crew later in the 

mission (figure 14) shorn the panel missing, the fuel cells on the shelf above 

the oxygen shelf tilted, and the high gain antenna damaged. 

The resultant loss of oxygen made the fuel cells inoperatire, leaving the Chl 

with batteries normally used only during reentry as the sole power source and 

with only that oxygen contained in a surge tank and repressurization packages. 

The lunar module, therefore, became the only source of sufficient battery ponyer 

and oxygen to permit safe return of the crew to earth. 

FIQURE 14 
40 

-PRESSURE RISE BEGiNS 

rTEMPERATURE RISE BEGINS 

SUMXARY ANALYSIS OF THE ACCIDENT 

The Board determined that combustion in oxygen tank #2 led to failure of 

that tank, damage to osygen tank #1 or its lines or valves adjacent to tank $2, 

removal of the Bay 4 panel and, through the resultant loss of all three fuel cells, 

to the decision to abort the Apollo 13 mission. In the attempt to determine the cause of ignition in oxygen tank #2, the course of propagation of the combustion, 

the mode of tank failure, and the way in which subsequent damage occurred, the 

Board has carefully sifted through all available evidence and examined the 

results of nearly 100 special tests and analyses conducted by the Apollo organization and by or for the Board after the accident. 

Although tests and analyses are continuing, sufficient information is now available to proride a clear picture of the nature of the accident and the events 

which led up to it. It is now apparent that the extended heater operation at 

KSC damaged the insulation on wiring in the tank and that this set the stage 

for the electrical short circuits which initiated combustion within the tank. 

While the exact point of initiation of combustion and the specific propagation 

path involved may never be known with certainty, the nature of the occurrence is sufficiently well understood to permit taking corrective steps to prevent its recurrence. 

The Board has identified the most probable failure mode. 

The following discussion treats the accident in its key phases : initiation, propagation and energy release, loss of oxygen tank So. 2 system integrity, and loss 

of oxygen tank No. 1 system integrity. Figure 15 shows the key events in the 

sequence. 

ACCIDENT EVENTS 

-HIGHEST PRESSURE READING 

-PANEL SEPARATES, TELEMETRY DROPS OUT 

-CREV~ REPORTS PROBLEM 

FUEL CELL 

NO. lFAlLS Ir 

tt 

I I I 

5553 55: 54 55:55 55:56 55:57 55:58 

GROUND ELAPSED TIME, HR:MIN 

FIGURE 15 

Initiation 

The evidence points strongly to an electrical short circuit with arcing as the initiating erent. Sear the end of the 65th hour of flight, about 2.7 seconds 

after the fans were turned on in the SJl oxygen tanks, an 11.1 ampere current 

spike and simultaneously a voltage drop spike were recorded in the spacecraft 

electrical system. Immediately thereafter current drawn from the fuel cells 
41 

decreased by an amount consistent with the loss of power to one fan. No other 

changes in spacecraft power were being made at the time. No power was on the heaters in the tanks at the time and the quaiitity gage and temperature sensor are rery lair. power clevices. The next anomalous event recorded was the beginning of a pressure rise in oxygen tank KO. 2, 13 seconds later. Such a time lag is possible with low level combustion at the time. These facts point to the 

likelihood that an electrical short circuit with arcing occurred in the fan motor or its leads to initiate the accident sequence. The energy available from the 

short circuit is estimated to hare been at least 10 to 20 joules. Tests conducted 

during this investigation hare shown that this energy is more than adequate 

to ignite Teflon wire insulation of the type contained within the tank. 

This likelihood of electrical initiation is enhanced by the high probability 

that the electrical wires within the tank were damaged during the abnormal 

detanking operation at KSC prior to lcunch. The likelihood of damage and the 

possibility of electrical ignition have been rerified by tests. 

Propagation 

While there is enough electrical power in the tank to cause ignition in the 

event of an arcing short circuit in defective wire, there is not sufficient electric 

power to account for all of the energy required to produce the observed pressure 

rise. 

There are materials within the tank that can, if ignited in the presence 

of supercritical oxygen, react chemically with the oxygen in heat-producing 

chemical reactions. The most readily reactire is Teflon, used for electrical insulation in the tank. Also potentially reactive are aluminum and solder. Our analyses indicate that there is more than sufficient Teflon in the tank, if reacted with 

oxygen, to account for the pressure and temperature increases recorded. Furthermore, the pressure rise took place over a period of more than 69 seconds, a relatively long period, and one which mould be more likely characteristic of 

Teflon combustion than metal-oxygen reactions. 

Thus, the Board concluded that combustion caused the pressure and temperature increases recorded in oxygen tank #2. The pressure reading for oxygen 

tank #2 began to increase abmit 13 seconds after bhe first electrical spike and 

about 55 seconds later the temperature began to increase. The temperature sensor reads local temperature. which need not represent bulk fluid temperature. Since 

the rate of pressure rise in the tank indicates a relatively slom propagation of 

burning along the wiring, it is likely that the region imlmediately around the temperature sensor did not become heated until this time. 

The data on the combustion of Teflon in supercritical oxygen in zero gwvity, 

developed in special tests in support of the Board, indicate that the rate of 

combustion is genecally consistent with these observations. 

Loss of oxygen tank #2 system integrity 

After the relatively slow propagation process described above took place, there was a relatively abrupt loss of oxygen tank #2 integrity. About 69 seconds after 

the pressure began to rise, it reached the peak recorded, 1008 psia, the pressure at 

which the cryogenic oxygen tank relief valve is designed to be fully open. Pres- sure began a decrease for 8 seconds, dropping to 996 psia before readings were 

lost. About 1.86 lsecoads after the last presumably valid reading from within the 

tank (a temperature reading) and .8 seconds after the last presumably valid pres- sure reading (which may or may not reflect the pressure within the tank itself 

since the pressure transducer is about 20 feet of tubing length distant), virtually 

all signal from the spacecraft was lost. Abnormal spacecraft accelerations were 

recorded apppoximately .42 seconds after bhe last pressure reading and approxi- mately .38 seconds before the 105s of signal. These facts all point to a relatively 

sudden loss of integrity. At about this time, ‘several solenoid valves, including 

the oxygen valves feeding two of the three fuel cells, were shocked to the 

closed position. The “bang” reported by the crew also occurred in this time period. 

Telemetry signals from Apollo 13 were lost for a period of 1.8 seconds. When 

signal --as reacquired, all instrument indicators from oxygen tank #2 were offscale, high or low. Temperatures recorded by sensors in several different locations in the service module showed slight increases in the several seconds folloiring reacquisition of signal. 

Data are not adequate to determine precisely the way in \rhich the oxygen tank #2 system failed. However, available information, analyses, and tests 

performed during this investigation indicate that the combustion within the 

pressure vessel ultimately led to localized heating and failure at the pressure 
42 

vessel closure. It is at this point, the upper end of the quantity probe, that 

the %-inch Inconel conduit is located, through which the Teflon insulated wires 

enter the pressure vessel. It is likely that the combustion progressed along the 

wire insulation and reached this looation where all of the wires come together. 

This, possibly augmented by ignition of other Teflon parts and even metal in 

the upper end of the probe, led to weakening and failure of the closure or the 

conduit or both. 

Failure at this point would release the nearly-1000 psi pressure in the tank 

into tlie tank dome, which is equipped with a rupture disc rated at 73 psi. 

Rupture of this disc or of tlie entire donie would then release oxygen, accompanied by combustion products, into Bay 4. The accelerations recorded mere probably caused by this release. 

Release of the oxygen then began to rapidly pressurize tlie oxygen shelf 

space of Bay 4. If tlie hole formed in the preswre vessel were large enough 

and formed rapidly enough, the escaping oxygen alone would be adequate to 

blow off the Bay 4 panel. However, it is also quite possible that the escape of 

oxygen ivas accompanied by comburtion of Mylar and Kapton (used extensively as thermal inwlation in the oxygen shelf compartment and in the tank dome) 

which n-ould augment the pressure caused by the oxygen itself. The slight 

temperature increases recorded at various locations in the service module indicate that combustion external to the tank probably took place. The ejected Bay 4 

panel then struck the high gain antenna, disrupting communications from tlie 

spacecraft for the 1.8 seconds. 

Loss of oruqen tank #1 integrity 

There is no clear evidence of abnormal behavior associated with oxygen 

tank #1 prior to loss of signal, although the one data bit (4 psi) drop in 

pressure in the last tank #1 pressure reading prior to loss of signal may 

indicate that a problem was beginning. Immediately after signal strength was regained, data show that the tank #1 system had lost its integrity. Pressure 

decreases were recorded over a period of approximately 130 minutes, indicating 

that a relatively slow leak had developed in the tank #1 system. Analysis has 

indicated that the leak rate is less than that which would result from a completely 

ruptured line, but could be consistent with a partial line rupture or a leaking 

check valve or relief valve. 

Since there is no evidence that there were any anomalous conditions arising 

within oxygen tank #1, it is presumed that the loss of oxygen tank #1 integrity 

resulted from the oxygen tank #2 system failure. The relatively sudden. and 

possibly violent, event asswiated with the failure of the oxygen tank #2 system 

could have ruptured a lir,e to oxygen tank #1, or have caused a valve to leak 

because of mechanical shock. 

APOLLO 13 RECOVERY 

lJnderst@nd ing tli c pro b lent 

In the period immediately following the Caution and Warning Alarm for Main 

Rus B undervoltage, and the associated “bang” reported by the crew, the cause of the difficulty and the degree of its seriousness were not apparent. 

The 1.8-second 105s of telemetered data was accompanied by the switching of 

the CSM high gain antenna mounted on the SJl adjacent to Bay 4 from narrow I)eam width to wide beam width. The high gain antenna (HGA) does this autornatically 200 milliseconds after its directional lock on the ground signal has been 

lost. 

A confusing factor was the repeated firings of various SM attitude control 

thrusters during the priod after data loss. In all probability. these thrusters were being fired to overcome the effects that oxygen venting and panel blow-off mere having on spacecraft attitude, but it was believed for a time that perhaps 

the thrusters n*ere malfunctioning. 

The failure of oxygen tank #2 and consequent removal of the Bay 4 panel 

produced a Phock which closed valves in the oxygen supply lines to fuel cells 

1 and 3. These fuel cells ceased to proride power in about three minutes, when 

the supply of oxygen between the closed valves and the cells was depletecl. 

The crew Tvas not alerted to closure of the oxygen feed valres to fuel cells 

1 and 3 because the valve position indicators in the CAI were arranged to gi7-e 

\varning only if both the oxygen and hydrogen valves closed. The hydrogen valves 

remained open. The crew had not been alerted to the oxygen tank #2 pressure 

rise or to its subsequent drop because a hydrogen tank lorn pressure \Yarning 
43 

llad blocked the cvogenic subsystem portion of the Caution and Tl‘arning systen1 several minutes before the accident. A limit seme light presumably came on in Jlission Control during the brief period of tank overpressure, but was not 

noticed. 

When the crelv heard the ‘,bang” and got the master alarm for low DC Main 

Bus B TFolbage, Love11 was in tlie lower equipment bay of the command module, 

stoJvil1g a televisio~l camera ivliich had just been in use. Hake was in the tunnel 

bet~veen the CSJI and the LJI, retunling to the CSJI. Swigert was in the left 

hand couch, monitoring spaceenaft performance. Because of the master alarm 

indicating lo^ voltage, Swigert moved across to the riglit hand couch where 

CSJI roltages can be observed He reported that voltages were “looking good” at 

5.3 :56 :lo. .it this time, voltage on Jlain Bus R had returned to normal levels and 

fuel cell5 1 and 3 did not fail for another lyg to 2 minutes. He also reported fluctuations in the oxygen tank #2 quantity, followed by a return to the off-scale 

high position. 

T7’11eii fuel cells 1 and 3 electrical output readings went to zero, the ground 

controllers coulcl not be certain that the cells had not somehow been disconnected 

froni their reispeetire buses and were not otherwise all riglit. Consequently about 

five minutes after the accident, controllers asked the crew to connect fuel cell 3 

to DC Main Bus B in order to be sure that the configuration was known. When it 

jj-as realized that fuel cells 1 and 3 were not functioning, tlie crew was directed 

to perform an emergency powerdown to reduce the load on the remaining fuel 

cell. Observing the rapid decay in oxygen tank #1 pressure, conbrollers asked 

the crew to re-power instrumentation in oxygen tank #2. When this was done. and 

it was realized that oxygen tank #2 had failed. the extreme seriousness of 4he 

situation became clear. 

During the succeeding periad, efforts were made to save the wmaining oxygen 

in the oxygen tank #1. Several attempti were made, bnt had no effect. The pressure. continued to decrease. 

It was obvious by about one-and-one-half liours after the aceidenit that the 

oxygen tank #1 leak coulcl not be stopped and that it would soon become necessary 

to use the LJI as a “lifeboat” for the remainder of the mission. 

By 68 :10, the 1,JI had been activated, the inertial guidance reference transferred from the CSJl guidance system to the LJI guidance yvstem, and the CSM 

systemcr were turned off. 

Return to carth 

The remainder of the mission was characterized by two main activitiesplanning and conducting the necessary prrrpulsion inaneurem to return the spacecraft to earth, and managing the use of consumables in such a way that the LM, 

which is designed for a basic mission with two crewmen for a relatively short 

duration, could support three men and serve as the control vehicle for the time 

required. 

One significant anomaly was noted during the remainder of the mission. At 

about 97 hours 14 minutes into the mission. Haise reported hearing a “thump” 

and observing venting from the LJI. Subsequent data review shows that the LJl 

electrical power system experienced a brief but major abnormal current flow 

at that time. There is no evidence that this anomaly was related to the accident. 

Analysis by the dpollo organization is continuing. 

.I number of propulsion options were clereloped and considered. It was neceqsary to return the spacecraft to a free-return trajectory and to make any re- quired midcourse corrections. Sormally, the Service Propulsion Systems ( SPS ) 

in the SM would be used for such maneuvers. However, because of the high 

electrical power requirements for using that engine, and in view of its uncertain 

condition and the uncertain nature of the structure of the SJI after the accident, 

it was decided to use the LJI descent engine if possible. 

The minimum practical return time was 133 hours to the Atlantic Ocean, 

and the maximum was 1.52 hours to the Indian Ocean. Recovery forces mere cleplo7ed in the Pacific. The return path selected was for splashdown in the 

Pacific Ocean at 142 90 g.e.t. This required a minimum of two burns of the 1,JI 

descent engine. A third hurn was subsequently made to correct the normal maneuver execution variations in the first two burns. One small velocity adjustment was also made with reaction control system thrusters. All burns were satisfactory. Figures 16 and 17 depict the flight plan followed from the tlme 

of the accident to splashdown. 
44 

MCC-7- 

x 

G. E.T. 142:54 

FIGURE 17 

The most critical consumables were water, used to cool the CSM and LM systems during use; CSM and LBf battery power, Itlie CSM batteries being for use during reentry and the LM batteries needed for the rest of the mission; LJl 

oxygen for breathing; and lithium hydroxide (LiOH) filter cannisters used to remove carbon dioxide from the spacecraft cabin atmosphere. These consumables, 

and in particular the water and LiOH cannisters, appear to be extremely marginal 

in quality shortly after the accident, but once the LM was powered down to 

conserve electric power and to generate less heat and thus use less water. 

the situation greatly improved. Engineers alt MCS developed a method which 

allowed the crew to use materials onboard to fashion a device allowing the use of the CM LiOH cannisters in the LBl cabin atmosphere cleaning system. At 

splashdown time, many hours of each consumable remained available. 

With respect to the steps taken after the accident, Mission Control and the 

crew worked, under trying circumstances, as well as was humanly possible, 

which was very well indeed. 

The Board's conclusion that the Apollo 13 accident resulted from an unusual 

combination of mistakes, coupled with a somewhat deficient and unforgiving 

design, is based on the Board's in-depth analysis of the oxygen tank, its design, manufacturing, test, handling, checkout, use, failure mode, and eventual effects on the rest of the spacecraft. 
45 

OXYGEN TANK # 2 HISTORY 

On February 26, 1966, the North American Aviation Corporation, now North 

American RocB\vell (NR) , prime contractor for the Apollo command and 

service modules (OSM), awarded a subcontract to the Beech Aircraft Corporation (Beech) to design, develop, fabricate, assemble, test, and deliver the Block 

I1 Apollo cryogenic gas storage subsystem. This was a follow-on to an earlier 

subcontract under which the somewhat different Block I subsystem was procured. 

4fanuf acture 

The manufacture of oxygen tank #2 began in 1966. In its review, the Board 

noted that the design inherently requires during assembly a substantial amount 

of wire movement inside the tank, where movement cannot be readily observed, 

and where possible damage to wire insulation by scraping or flexing cannot be 

easily deteated before the tank is capped off and welded closed. It does not 

appear, however, that these design deficiencies played any part in the accident. 

Several minor manufacturing fiaws were discovered in the oxygen tank #2 

in the course of testing. A porosity in a weld on the lower half of the outer shell 

necessitated grinding and rewelding. Rewelding was also required when it was determined that incorrect welding wire had been inadvertently used for a small 

weld on a vacuum pump mounted on the outside tank dome. The upper fan motor 

originally installed was noisy and drew excessire current. The tank mas disassembled and the heater assembly fans, and heaters were replaced. 

Following acceptance testing at Beech ; during which the tank was filled and 

detanked without apparent difficulty, oxygen tank #2 was shipped to NR on May 3, 1967, for installation, which was completed on March 11, 1968, on a shelf 

to be installed in service module 106 for flight in the Apollo 10 mission. 

From April 27 to May 29, 1968, the assembled oxygen shelf underwent standard proof pressure, leak, and functional checks. One valve on the shelf leaked 

and was repaired, but no anomalies were noted with regard to oxygen tank #2, 

and therefore no rework of oxygen tank #2 was required. 

On June 4, 1968, the shelf was installed in SM 106. 

Between August 3 and August 8, 1968, testing of the shelf in the SM was conducted, including operation of the heater controls and fan motors. So 

anomalies were noted. 

Due to electromagnetic interference problems with the vacuum pumps on cryogenic tank domes in earlier Apollo spacecraft, a modification mas introduced 

and a decision mas made to replace the complete oxygen shelf in SRI 106. An 

oxygen shelf mas approved modifications was prepared for installation in SM 

106. On Odober 21, 1968, the oxygen shelf was removed from SlI 106 for the 

required modification and installation in a later spacecraft. 

During the initial attempt to remove the shelf, one shelf bolt mas mistakenly 

left in place ; and as a consequence, after the shelf was raised about two inches, 

the lifting support broke, allowing the shelf to drop back into place. At the 

time, it was believed that the oxygen shelf had simply dropped back into place, 

and an analysis was performed to calculate the forces resulting from a drop of 

two inches. It now seems likely that the shelf was first accelerated upward and 

then dropped. 

The remaining bolt was then removed, the incident recorded, and the osygen 

shelf was removed without further difficulty. Following removal, the osygeii shelf mas retested to check shelf integrity, including proof pressure tests, leak 

tests, and fan and heater operation. T’isual inspection revealed no problem. 

These tests would have disclosed external leakage or serious internal malfunctions of most types, but would not disclose fill line leakage within oxygen tank 

#2. Further calculations and tests conducted during this investigation have 

indicated that the forces experienced by the shelf were probably close to those 

originally calculated, assunling a 2-inch drop only. The probability of tank damage from this incident, therefore, is now considered to be rather low, although 

it is possible that a loosely fitting fill tube assembly c.ould have been displaced 

by the event. 

The shelf passed these tests and was installed in SJI 109, the dpollo 13 serv- ice module, on Sovember 22. 1068. The shelf tests accomplished earlier in SJI 

106 were repeated in RJI 109 in late December and early January, with no significant problems, and SJI 109 was shipped to KSC in *June of 1969 for further 

testing, assemblg on the launch vehicle, and launch. 

474x6 0-70-4 
46 

Testing at RNC 

At the Kennedy Space Center the CM and the SM were mated, checked, as- sembled on the Saturn 1‘ launch vehicle, and the total vehicle was moved to the 

launch pad. 

The Countdown Demonstration Test (CDDT) began on Narch 16, 1970. Up 

to this point, nothing unusual about oxygen tank #.2 had been noted during the 

extensive testing at KSC. Cryogenic oxygen loading and tank pressurization to 

33 psi was completed without abnormalities. At the time during CDDT when 

the oxygen tanks are normally rented down to about .50 percent of capacity, 

oxygen tank #1 behaved normally, but oxygen tank #2 only went down to 92 

percent of its capacity. The normal procedure during CDDT to reduce the 

quantity in the tank is to apply gaseous oxygen at 80 psi through the vent line 

and to open the fill line. When this procedure failed, it was decided to proceed 

with the CDDT until completion and then look at the oxygen detanking problem 

in detail. 

On Friday, March 27, 1970, detanking operations were resumed. after discussions of the problem had been held with KSC, NSC, SR, and Beech personnel 

participating, either personally or by telephone. As a first step, oxygen tank #2, 

which had self-pressurized to 178 psi and was about 83 percent full. was vented 

through its fill line. The quantity decreased to 65 percent. Further discussions 

between KSC, MSC, KR, and Beech personnel considered that the problem might 

be due to a leak in the path between the fill line and the quantity probe due to 

loose fit in the sleei-es and tube. Such a leak would allow the gaseous oxygen 

being supplied to the rent line to leak directly to the fill line without forcing 

any significant amount of LOX out of the tank. At this point, a Discrepancy 

Report against the spacecraft system mas written. 

A “normal” detanking procedure was then conducted on both oxygen tanks, 

pressurizing through the vent line and opening the fill lines. Tank #1 emptied in 

a few minutes; tank #2 did not. Additional attempts were made with higher 

pressures without effect, and a decision was made to try to “boil off” the remaining oxygen in tank #2 by use of the tank heaters. The heaters were energized 

with the 66 volt DC GSE power supply and, about 1% hours later, the fans were turned on to add more heat and mixing. After 6 hours of heater operation. the 

quantity had only decreased to 35 percent, and it was decided to attempt a pressure cycling technique. With the heaters and fans still energized, the tank 

lvas pressurized to about 300 psi, held for a few minutes, and then vented through 

the fill line. The first cycle produced a 7 percent quantity decrease, and the 

process was continued, with the tank emptied after five pressure/vent cycles. 

The fans and heaters were turned off after 8 hours of heater operation. 

Suspecting the loosely fitting fill line connection to the quantity probe inner 

cylinder, KSC personnel consulted with cognizant personnel at MSC and at SR. 

It was decided that if the tank could be filled, the leak in the fill line would not 

be a problem in flight, since it was felt that even a loose tube resulting in an electrical short betlveen the capacitance plates of the quantity gage woulid result 

in an energy level too low to cause any other damage. Replacement of the oxygen 

shelf in the CM would have been difficult and mould have taken at least 45 hours. 

In addition, shelf replacement would have had the potential of damaging or de 

grading other elements of the service module in the course of replacement activity. 

Therefore, the decision was made to test the ability to fill oxygen tank #2 on March 30, 1970, 12 days prior to the scheduled Saturday, April 11, launch. so as to be in a position to decide on shelf replacement well before the launch 

date. 

Flow tests were first made with gaseous oxygen on oxygen tank #2 and on oxygen tank #1 for comparison. No problems were encountered. and the flow 

rates in the two tanks were similar. In addition, Beech was asked to test the 

electrical energy level reached in the event of a short circuit between plates of 

the quantity probe capacitance gage. This test showed that very lorn energy 

levels would result, Then, oxygen tanks #1 and #2 were filled with LOX to 

about 20 percent of capacity on March 30 with no difficulty. Tank #1 emptied in 

the normal manner, but emptying oxygen tank #2 again required pressure 

cycling with the heaters turned on. As the launch date approached, the oxygen tank #2 detanking problem was considered by the Apollo organization. At this point, the “shelf drop” incident 

on October 21, 1968, at SR was not considered and it was felt that the apparently 

normal detanking which had occurred in 1967 at Beech was not pertinent because it was believed that a different procedure was used by Beech. In fact, 
47 

however, the last portion of the procedure was quite similar, although at a slightly lower pressure. Throughout these considerations, which involved technical and managenleiit 

personnel of KSC, MSC, SR, Beech, and SASh Headquarters. emphasis was directed toward the possibility and consequence of a loose fill tube ; rery little attention was paid to the extended heater and fail operation. except to note that they 

opeiated during and after the detanking sequences. Many of the principals in the discussion were not aware of the extended 

heater operations. Those that did know the details of the procedure did not 

consider the possibility of damage due to excessive heat within the tanli, and 

therefore did not advise management officials of any possible consequences of 

the unusually long heater operations. 

As I noted earlier, each heater is protected with a thermostatic switch, niounted on the heater tube, which is intended to open the heater circuit when it senses a temperature of about 80°F. In tests conducted since the accident, however, 

it was found that the switches failed to open when the heaters were powered 

from a 63 volt DC supply similar to the power used at KSC during the detanking sequence. Subsequent investigations have shown that the thermostatic 

switches used, while rated as satisfactory for the 28 volt DC spacecraft power 

supply, could not open properly at 63 volts DC with 6-7 amps of current. A 

review of the voltage recordings made during the detanking at KSC indicates 

that, in fact, the switches did not open when the temperature of the switches rose past 80" F. Figure 18 shows a thermostatic switch welded closed after ap- plication of ly~ amperes of f% volts DC. Further tests have shown that the temperatures on the heater tube subsequent to the switch failures niay have reached as much as 1000°F during the detanking. This temperature can cause serious 

damage to adjacent Teflon insulation, and such damage almost certainly OCcurred. Figures 19 and 20 shorn the condition of mires, such as those used in 

the fan motor circuit, after they have been subjected to temperatures of about 

1000°F. 

FIGURE 18 
48 

FIGURE 19 

FIGURE 20 
49 

None of the above, however, was known at the time and, after extensive con. sideration was given to the possibilities of damage from a loose fill tube, it was 

decided to leave the oxygen shelf and oxygen tank #2 in the SJI and to proceed 

with preparations for the launch of Apollo 13. In fact, following the special detanking, the oxygen tank #2 mas in a hazardous condition whenever it con- tained oxygen and was electrically energized. This condition caused the Apollo 

13 accident, which was nearly catastrophic. Only the outstanding performance on the part of the crew, Mission Control, and other members of the team which 

supported the operations, successfully returned the crew to earth. 

In investigating the Apollo 13 accident, the Board attempted to identify those 

additional technical and management lessons which can be applied to help as- sure the success of future spaceflight missions. Several recommendations of this 

nature are included. 

RECOMMENDATIONS 

Before reading the Board’s recommendations, I would like to point out that 

each Member of the Board concurs in each finding, determination, and recom- mendation. 

The Board’s recommendations are as follows : 

1. The cryogenic oxygen storage system in the service module should be 

a. Remove from contact with the oxygen all wiring, and the unsealed 

motors, which can potentially short circuit and ignite adjacent materials ; or otherwise insure against a catastrophic electrically induced 

modified to : 

fire in the tank. 

b. Minimize the use of Teflon, aluminum, and other relatively com- bustible materials in the presence of the oxygen and potential ignition 

sources. 

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

operational problems. 

3. The warning systems onboard the Apollo spacecraft and in the Mission 

Control Center should be carefully reviewed and modified where appropriate, with specific attention to the following: a. Increasing the differential between master alarm trip levels and 

expected normal operating ranges to avoid unnecessary alarms. 

b. Changing the caution and n7arning system logic to prevent Ian outof-limits alarm from blocking another alarm when a second quantity 

in the same subsystem goes out of limits. c. Establishing a second level of limit sensing in Mission Control on critical quantities with a visual or audible alarm which cannot be easily 

overlooked. 

Id. Providing independent talkback indicators for each of the six fuel 

cell reactant valves plus a inaster alarm when any valve closes. 

4. Consumables and emergency equipment in the LM and the CJI should 

be reviewed to determine whether steps should be taken to enhance their 

potential for use in a “lifeboat” mode. 

3. The JIanned Spacecraft Center should complete the special tests and 

analyses now underway in order to undersband more completely the details of the Apollo 13 accident. In addition, the lunar module power system anomalies should receive careful attention. Other SASA Centers should 

continue their support to MSC in the areas of analysis and test. 

6. Whenever significant anomalies occur 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. Furthermore, critical 

decisions inrolving the flightworthiness of subsystems should require the 

presence and full participation of an expert who is intimately familiar 

with the details of that subsystem. 

7. SASA should conduct a thorough reexamination of all of its slyacecraft, launch vehicle, and ground systems which contain high-density oxygen. or other strong oxidizers, to identify and evaluate potential combustion 

hazards in the light of information developed in this investigation. 

8. SASA should conduct additional research on materials compatibility. 

ignition, and combustion in strong oxidizers at various g levels ; and on the 

characteristics of supercritical fluids. Where appropriate, new SAW.% design 

standards should be developed. 
50 

9. The Manned Spacecraft Center should reassess all Apollo spacecraft 

subsystems, and the engineering organizations responsible for then] at JISC 

and at its prime contractors, to insure adequate understanding and control 

of the engineering and manufacturing details of these subsystems 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. 

CONCLUSION 

In concluding, I would stress two points. 

The first is that in this statement I hare attempted to summarize the Board's 

Report. This Report and its appendices are the result of more than seren weeks 

of intensive Kork by the Board. its Panels. and staff. supported by the SASA 

and contractor organizations. In the interest of time. I hare not included manr supporting findings and determinations which are set fort11 in the Report. 

The second point I wish to make is this : The Apollo 13 accident. which aborted man's third mission to explore the sur- face of the moon, is a harsh reminder of the immense difficulty of this undertaking. 

The total Apollo system of ground complexes. launch rehicle, and spacecraft constitutes the most ambitious and demanding engineering development ever undertaken by man. For these missions to succeed, both men and equipment must 

perform to near perfection. That this system has already resulted in two suc- cessful lunar surface explorations is a tribute to those men and women who 

conceived. designed, built, and flew it. 

Perfection is not only difficult to achieve, but difficult to maintain. The imperfection in Apollo 13 constituted a near disaster, averted only by outstanding 

performance on the part of the cren- and the ground control team which supported them. The Board feels that the Apollo 13 accident holds important lessons which. 

when applied to future missions, will contribute to the safety and effectireness of 

manned space flight. 

JIr. Chairman, this concludes my prepared statement. 

STATEJIEXT BY DR. PAINE 

The CH.IIRifAhr. Dr. Paine, go ahead. 

Dr. PAINE. Mr. Chairman, inembers of the committee, in our appearance before the committee on April 24, 1970, Apollo Program 

Director Dr. Rocco Petrone, Flight Director Glynn Lunney, and 

astronauts Jim Love11 and Jack Swigert reported to you our understanding as of that time of the events leading to the accident 

and the subsequent operations which brought the astronauts safely 

back to earth. At the same hearing, I reported to you the actions Dr. 

Low and I had taken to assure a prompt, thorough, and objective investigation of the accident. These included : 

REVIEWS ACTIONS FOLLOWING .ICCIDENl' 

(1) The establishinent of the Apollo 13 Review Board. with JIr. 

Edgar 31. Cortright, Director of the Langley Research Center, as 

Chnirinan. 

(2) The instruction to K'SSd's Aerospace Safety Advisory Panel 

to review the procedures and findings of the Apollo 13 Review Board 

and to submit its independent report within 10 days of the Review 

Board's report, and 

(3) The instruction to Dr. Dale Myers, NASA's Associate hdministrator for Manned Space Flight, to provide necessary support to 

the Apollo 18 Review Board and to make recommendations, also with- 
51 

in 10 days of the Review Board's report, on plans for eliminatillg 

tlie problems encountered in Apollo 13 in order to proceed with 

,A.pollo 14 and future manned space missions. 

REVIEWS RESULTS OF ACTIONS 

Today Tve are here to review with you the results of these actions and the resulting future program actions which Dr. Lorn and 

I am iio\v taking to preclude a recurrence of such accidents and to 

111ove ahead with the Nation's manned space flight program. In summary : 

The report of the Apollo 13 Review Board was presented to us by 

Xr. Cortright on June 15 and made available to the committee on the 

same day. Dr. Lov and I have now had an opportunity to study 

the report in detail and to reyiew carefully its recommendations. In 

our vie\T- it is an excellent report based 011 a thorough and objective 

investigation and highly competent analysis. It clearly pinpoints the 

causes of the Apollo 13 accident and sets forth a comprehensire set, 

of recommendations to guide our efforts to prevent the occurrence of 

similar accidents in the future. 

The Aerospace Safety Advisory Panel submitted its report to NASA 

inanagement at a meeting in TTrasliiiigton June 25, 1970. With your 

permission, Mr. Chairman, I would like to place its report which 

is in the forin of a letter from its Chairman, Dr. Charles D. Harrinpton, in the record. 

The CHAIRMAN. Without objection, that will be done. 

(For the letter referred to see p. 56.) 

Dr. PAISE. Dr. Harriiigton is here this morning to respond to any 

questions you may have. At this point I would like to read the key 

portions of his letter report summarizing the Safety Panel's appraisal 

of the job done by the Apollo 13 Review Board. He says : 

The Panel found that the Board' prwedures and scope of inquiry proved effective in their task. The Review Board has performed a thorough and technically 

competent analysis in the reconstruction of the factors contributing to the Apollo 

13 abort. We found no evidence and no reason to doubt the technical ralidity of 

their deterinination and findings. 

This independent evaluation provides substantial additional confidence to Dr. Low and to me that our farorable appraisal of the report is correct. 

Dr. Myers, Dr. Petrone, and the Office of Manned Space Flight 

have also completed extensive experiments, tests, studies, reviem, redesign work and program rescheduling activities, and have presented 

recommendations on the required correctire measures and program 

adaptations. Last Thursday Dr. IAW and I held an extensive revielv at 

which Dr. Myers, Dr. Petrone, Colonel McDivitt and other officials 

of the Apollo program discussed in detail the technical problems and 

alternatives \l+h the senior officials of NhSh. Also present \yere 

Mr. Cortright and members of the Reviev Board. Dr. Harringtoll and 

members of the Aerospace Safety Advisory Panel, hlr. TITilliam A. 

&knders, Executive Secretary of the National Aeronautics and Space 

Council, and the Directors of N,4SA's Manned Space Flight centers : 

Dr. Robert R. Gilruth, Dr. Kurt H. Debus, and Dr. Eberhard Rees. 

Based on the discussions at this review and at follo\vup meetings ex- 
52 

tending oTTer the next 2 days, Dr. Myers has formally submitted to me 

with his endorsement the filial recommendations of Dr. Petrone, the 

Apollo Program Director. These are embodied in Dr. Petrone‘s memorandum to me of June 27, 1970, which has been made available to the 

committee, and which I would like to place in this record, with your 

permission, Mr. Chairman. 

The CHAIRXAN. TT’ithout objection : that will be done. 

(For the ineinorandum referred to see p. 57.) 

Dr. PAISE. On the basis of the reports and recoiiiiiiellclatiolis before 

us and detailed discussions with responsible and knowledgeable experts 

in SdSA, Dr. Low and I have approved the following actions to 

implement the recoininendations of the Apollo 13 Review Board and to 

carry out the steps recoininended by Dr. Petrone and Dr. JIrers to prepare for the Apollo 14 mission. In summary, these actions are: 

APOLLO 14 POSTPOKED TO EARLY 1971 

First, the recoinmendations of the Apllo 13 ReT-iem Board will be 

implemented before the Apollo 14 mission is approved for laulich. 

This will require postponing the launch date to no earlier than January 31,1971. Command Service Module systems will be modified along 

the recommended lines to eliminate potential combustion hazards in 

high pressure oxygen of the type revealed by the Apollo 13 accident. 

Tlnsealed fan motors will be removed from the oxygen tanks and an 

additional oxygen tank added to the service module of ,kpollo 14. 

Electrical wiring within high pressure oxygen systems which might 

provide an ignition spark if damaged will be limited to stainless steel 

sheathed wires. Teflon, aluminum, and other potentially reactiT-e materials in the presence of high pressure oxygen will be used as little as 

possible and kept away from possible ignition sources. For example, 

the quantity probe will be stainless steel instead of aluininnm and the 

fuel cell oxygen supply valve which now has Teflon-insulated wires in 

high pressure oxygen will be redesigned to eliminate this hazard. 

Warning svstems on board the spacecraft and at mission control 

mill be modified consistent with the Board‘s recommendations to provide more immediate and risible warnings of system anomalies. *4 

comprehensive reviev of spacecraft emergency equipment and procedures and use of command service modules and lunar modules in 

“lifeboat” modes is now underway at the Manned Spacecraft Center in Houston. Dr. Petrone will outline for you the specific actions 

we plan to take in response to the first six recommendations of the 

Board, and Dr. Myers will discuss his specific plans for critically reassessing all Apollo spacecraft subsystems in response to recommendation No. 9 of the Board. 

NASA REVIEW OF APOLLO 13 REPORT 

Secondly, the associate administrators in charge of the Offices of 

Space Science and Applications, Manned Space Flight, and Advanced 

Research and Technology, have been directed to review the Apollo 13 

review board report to apply thronghont NASA the lessons learned 

in their areas of responsibility. They have been instructed to take 

action with respect to recommendation KO, 6 (concerning anomalies 
53 

in c,ritical subsystems prior to flight), recommendation No. 7 (calling 

for a thorough reexamination of all spacecraft, laiinch vehicle and 

ground systems which contain strong oxidizers to evaluate potential 

hazards) and recommendation No. 9 (concerning the design, manufacture, test, and operation of spacecraft subsystems). I have requested a written report by August 25 on their assessment and the 

actions taken or proposed. 

In addition, we will take steps to disseminate widely throughout the 

industry and the technical community the lessons of Apollo 13 to prevent recurrences in other areas. You might be interested to know in 

this connection that I have forwarded to Academician Keldysh of the 

Soviet Scademy of Sciences a copy of the complete Apollo 13 Review 

Board report so that lessons which might be learned from our accident 

can be applied to prevent a similar hazard to Soviet cosmonauts. 

Third, the Aerospace Safety Research and Data Institute (*ISRDI) 

at the K'BSA Lewis Research Center has been directed to conduct 

additional research on materials compatibility, ignition, and combustion at various gravity levels, and on the characteristics of supercritical fluids, as recommended by the A4p~llo 13 Review Board. This 

will expand a review already begun by ASRDI on oxygen handling 

in aerospace programs. In this effort, the Lewis Research Center vill . be supported by other elements of the NASA organization. This 

research will be of direct long-term benefit to NASA in carrying out 

its future programs, and will help other sectors of the economy. 

-\EROSPACE SAFJ3TY .IDVISORI' PAXEL 

Fourth, I have requested that the Aerospace Safety hd~isorg Panel 

conduct a review of the management jirocesses utilized by XASA in 

im lementing the recommendations of the A4p0110 13 Review Board 

an$ report to me their views no later than the Ap0110 14 flight readiness review. This will again gh-e us the benefit of the panel's valuable 

independent insight when future decisions are made. I have also asked 

Mr. Cortright to reconvene the A4pollo 13 Review Board later this year, 

as he suggested, to review the results of continuing tests to determine 

whether any modifications to the board's findinp, determinaltions, or 

recommendations are necessary in light of additional evidence which 

may become available. 

JANUARY LIUSCH OF APOLLO 14 POSSIBLE 

The assessment of the Office of Manned Space Flight, in which 

Dr. Low and I concur, is that the reasonable time required for the 

design, fabrication, and qualification testing of the modifications to 

the Apollo system we have determined to be necessary, and for the 

other actions outlined above which must be taken before the next 

hpollo mission, Id1 permit us to launch Apollo 14 to the Fra Mauro 

region of the moon at the January 81, 1971, launch opportunity. This 

will also move the plaiined launch date for Apollo 16 several months 

to July or August 1971, maintaining the approximate 6-month interval 

between launches on which our operations in the Apollo program 

are now based. However, we will not launch Apollo 14 or any other 

flight iinless and until we are confident that we have done ererything necessary to eliminate the conditions that caused or contributed 
54 

to tlie problems we encountered on A4pollo 13 and are re:icly in all 

other respects. One of our prime coiicerns will be to maintain the 

efficiency aiid high standard of perfoiwinnce requiiwl of our Ixmicli 

aiid ground snpport teams dnrinp the extended periods of reclacecl 

activity entailed by the revised mission schedule niid by the substantial cntbacks which have been made necessai*y by the overall rediictioiis in tlie Kaition's space progrnni. 

JIODIFICATION COSTS ESTIXiTED AT $1 S JIILLION 

It is too early to present to you our detailed estimates of tlie costs 

and budgetary impact of the spacecraft modifications and program 

changes that we are now making. Our best current estimate is that the 

modifications aiid changes related to the actions resulting froin the 

Apollo 13 accident will be in the range of $10 to $15 million of increased costs, which we plan to handle within our total hpollo buclget. 

Before turning to Dr. Myers and nr. Petrone, I ~vould like to coniinent briefly on tlie lessons to be learned froin Apollo 13. The Review 

Board found "that tlie accident was not tlie result of a chance iiialfunction in a statistical sense. but rather resulted froin an unusual conibination of mistakes, coupled with a soniewliat deficient and unforpiving 

design." The presence of inadeqnate thermostatic switches in the 

heater circuits of the oxygen tanks. the loose fill tube assembly probably caused by a buildup of "worst case" tolerances and the "shelf 

dropping" incident, the iml3rovised det ailking procedure employed 

in preparing for launch, and the resnlting dainaped Teflon-insulated 

fan motor wiring caused by overheating ~vliich later provided the 

ignition spark-together all of these elenieiits combined to cause the 

accident. In the absence of any one of these links in the chain of events. 

oxygen bottle No. 2 would not have failed. 

NASS's actions in response to the Roqrd's recommendations will, 

in my TTiew. avoid those specific things which led or contributed to the 

Apollo 13 accident : and the reviews and research we hare uiidertaken 

will help us aroid future potential hazards throughout our programs. 

Rut in a larger context, we at NLiSA must be coiiceriied with the fact 

that despite the rigorous management controls in effect and, from all 

the evidence. adhered to, a hazardous condition existed that was not 

identified aiid corrected. In fact, the presence of the inadequate thermostatic switches in the tank and tlie resultant baking of the wires at 

temperatures as high as 1000" F. during detankinp were not discovered 

until actual full-scale tests were conducted for the Review Board in 

vhich wires were damaged. leading to ii reexamination of the data 

recorded at Kennedy Space Center during the detaiikiiig and the 

switch specific a t' 1011s. 

REVIEW OF BI'.\CECRAFT COXTR.-\CTS 

With regard to our contracts ivitli Sorth American Rock~vell and 

Gruminnii for the spacecraft inr-olved in the hpollo 13 mission, we 

hare underway a review of tlie incentive provisions in their contracts 

to determine what steps should be taken by NASA in light of the 

accident. In accordance with our contract ivith Sortli hinericaii Rockwell we will take the sen-ice module oxygen system failure into accoullt 
55 

in determining the amount of tlie 1970 award fees to be paid. That 

fee will be determined in view of all activities during 1070, and thus 

will be based not only on the Apollo 13 accident but also on the effectiveness of the redesign and rebuilding activities duriiig the inoiiths 

following tlie accident in preparation for Apollo 14. 

In the case of the Griimman lunar iiiodule contract the fee provisions 

are phrased only in terms of performance during an actual lmmr landing mission. However, since in performing as it did in tlie "lifeboat" 

mode the lunar module "8qiiarius" clearly demonstrated its ability to 

hnve successfully performed inost of the operatioiis of an actnal landing, we are performing a technical assessment of the Apollo 18 inissioii 

as it was flo~n to establish what portion of the performance was 

deiiioiistrated and, therefore, what portion of the incentive fee should 

be paid now. 

In a program as large and complex as Apollo, iiivolviiig thonsands 

of people throughout tlie country, we must obriously depend 011 a 

rigorous docuiiieiitatioii system to record and convey program maiiageiiient information, TTliat we must always guard against. however, is 

tlie possibility of permitting this flow of careful docuiiieiitation to 

substitute for the meaningfnl excliaiige of information. No matter I~OV 

thorough aiid carefnl Ire are, we nltiinately depend on incisive and 

informed problem analysis by competent people who make the key 

decisioiis on the basis of their thorougli understanding of tlie underlyiiig actualities which are recorded in the documentation. 

We caiiiiot in tlie case of Apollo 13, point to one iiidiridual or group 

of individuals or organization aiid say that they caused the accident. 

Nor hare ire 01' tlie Review Board been able to formulate-even with 

all the advantages of hindsight-a management procedure which, had 

it been in effect for Apollo 13, would hare guaranteed that such an 

accident could never liappen. The excellent recommendations of tlie 

Board in tlie areas of management aiid procedures caii further 

strengthen Apollo aiid other KASA programs. But in the last analysis, 

\re must depend npon tlie thoroughness mid detailed understanding of 

all those in 1-esponsible positions in the SASA-industry hierarchy 

throughout el-ery plinse of design, niannfacture, test aiid flight operations. I have the utmost confidence that the SASA team caii fix the 

Apollo 13 problem aiid strengthen its operations to minimize tlie 

cliaiices of futui*e problems. We realize, liowerer-and the members 

of this committee realize-that the exploration of space is a deinaiidiiig 

and hazardous enterprise in wliicli man is probing the unknoil-n. 

SASh men aiid women are doing iiiaiiy things for the first time. 

deficiencies in our ability to look ahead aiid foresee difficulties, any 

inattention to detail will be exposed in the harsh enrironments in which 

our v-ork is tested. In my opinion, no finer or more dedicated group of 

people lias ever worked together more effectively than this Satioii's 

space team, aiid I aiii confident of tlieir continuing future success. 

Mr. Chairman, this concludes my statement, and I would now 

like to ask J1r. Myers aiid Dr. Petroiie to summarize in detail for you 

our proposed actions in respoiise to the recommendations of the Apollo 

13 Review Board. 

(The letter and meiiiorandiiin referred to 011 page 51 are as 

follon-s :) 
56 

NATIOR’AL AEROKAUTICS AND SPACE ADMINISTRATION, 

OFFICE OF THE ADMIXISTRATOR, 

TVnshington, D.C., June 25, 1970. 

Dr. T. 0. PAIKE, 

Administrator, Nationnl Aeronautics awl Space Administration, 

Washington, D.C. 

DEAR DR. PAINE: This letter is in response to your request that the Aerospace Safety Advisory Panel review and comment on the procedures and findings of the Apollo 13 Reriew Board. 

With the backgorund of the Panel’s earlier review of the Apollo program 

management and risk assessment system, 7T-e met in working sess,ions with 

JIr. Cartright and his panel chairmen as well as v-ith niemkrs of the Ap~llo 

13 Inrestigation Team. Thus n-e hare been able to review the erolring 

process of inquiry and to obsen-e much of the evidence as it was dereloppd. 

Our participation in the Board’s deliberations proridecl a timely foruni for 

both their consideration of our comnients ancl our uncler,ctancling of their 

determinzation and findings. 

The Panel found that the Board’s procedures and scope of inquiry prored 

effectire in their task. The Review Board has perforined a thorough and 

technically competent analysis in the reconstruction of the factors contributing 

to the dpollo 13 abort. We found no evidence or suggestion that significant data or events were not pursued with diligence and no reason to doubt the techniml 

validity of their determination and findings. 

Recommendations made for the redesign of the system should improre the 

manufacturability and niinimize the failure modes of the system under abnormal conditions. The recomenclations for impro~einents in the analysis and 

validation of nonstandard situations are certainly necessary and should be 

implemented immediately. The efficacy of the other recomniendations for con- tinuing research, assessment and engineering organination review will depend 

upon the manner in which they are implemented. There must no^ be prorisionc 

for the performance of these recomniendations in a manner satisfactory and 

risible to you. 

While there is no need for us to file a separate lengthy report, we nould like 

to emphasize certain points. 

The total Apollo system, both hardware and software. possePses a considerable degree of inherent redundancy and reliability. Snccessful return to earth 

after a major system failure was possible hecause many alternate systems. 

modified procedures and non-standard operations were arailable to the ground 

and flight crew. Almost all of the swcial actions ancl procedures required for a successful 

recovery had been thought out and dereloped in the premission period. This 

says much for the thoroughneqs of mission planning. The mission control group responded to the unusual and critical wents in 

a very effective manner. Their high degree of skill and knowledge. and the 

discipline to make it effectire in a timely fashion. were as critical as the crew resnonse in achieving the recovery of Apollo 13. 

The elaborate technical management control system created in iinollo has, in our experience, achieved a high degree of maturity and definition. In this case. 

the acceptance of the design and manufacturahility risks, the lack of understanding of thermal switch performance and the informal system in the eralnation of the test procedure change reflect more a failure of human judgment 

than a failure in the requirements of the technical management system. 

Such an incident brings into prominence the question of management control 

ancl risk acceptability. This incident should not call into question the basic 

credihilitg of the technical management system. The system now in esistence 

has significantly reduced the possihility of errors in human judgment. The 

phasedown of personnel has lieen achieved in a manner to nlasimize skill retention. and the flights of prior Apollo Saturn Systems hare been exninently sue. 

cessful. The problem now is to sustain hiininn motivation and assure Torking familiarity mith the suhsystenis and their hazards. Finally erosion in engineering support and launch operations manpower and untimely delays in restllning 
57 

the launch schedule could only serve to degrade the capability to service and 

launch the Apollo Saturn System and thus introduce unknown risks. 

In summary, the Panel commends the integrity of effort of the Board and its 

associates and now awaits the implementation of its recommendations. 

Sincerely yours, CHARLES D. HARRINGTON, 

Ckaimtalz, Aerospace Safety Advisory Palzel. 

NATIONAL AEROKAUTICS AND SPACE ADMIXISTRATIOX, 

Washington, D.C., June 27,1970. 

To : A/Administrator. 

Through : N/hssociate Administrator for Manned Space Flight. 

From : JIA/Apollo Program Director. 

Subject : Proposed Actions in the Apollo Program in Response to the Apollo 13 

The OJISF response to the Apollo 13 Review Board Report mas presented to the 

ddlninistrator at a nieeting on June 25, 1970. Based on discussions which tool; 

place at that meeting, we have done additional analysis and testing. Results of 

these efforts, which hare been discussed at several meetings, have led me to 

recommend the following actions : 

(a) Replacement of all Teflon coated wires by stainless steel sheathed 

wires 

(b) Removal of unsealed fan motors 

(c) Jfodification of heater configuration from two 76-watt heaters 

powered from a single bus to three 50-watt heaters powered from two 

independent buses. 

(d) Jlodification of quantity probe from aluminum to stainless steel 

(e) Addition of temperature sensor to heater probe 

(f) Elimination of heater thermal switches 

(g) Xodiflcation of tank cap to preclude need for rotation of quantity 

2. Addition of third oxygen tank to avoid operation in loir quantity 

regime on hpollo 14 & 15 thereby permitting remoral of unsealed fan motors : and review the oxygen subsystem requirments for Apollo 16 and subsequent 

missions. 

3. Modification of fuel cell oxygen reactant rake to separate Teflon 

coated mires from oxygen environment. 

4. Caution and warning system modifications 

(a) Modification of caution and warning system to provide an audio 

(b) Installation of talkback indication for single fuel cell reactant 

(c) Installation of existing modification kit to adjust, where appro5. Addition of secand-level limits sensing in the JIission Control Center as reconieuded by the Review Board. 

6. Completion of the comprehensive review 110~- under way of the con sumaliles and emergency equipment in the Lunar JIoclule and Command 

Jlodule in accordance with recommendation #4 of the Revien- Board 

Report. 

'7. Conipletion of the special tests and analyses no~v under ~vay in accordance with recommendation #5 of the Review Board Repolrt. 

8. Respond prior to August 15th to your directions with regard to reco1iinlendation #6, 7, and 0 of the Board Report in so Ear as they apply to the 

Apollo Program. 

Based on the estimated time required for the design, fabricatioll and qnalification testing for the niodifications outlined above, and for the other actioni 

set forth, I propose the h1mllo 14 mission Iir sclieduled for lanllcll to the Fra 

Jlauro region of the moon no earlier than the 31 January 1971 launch opllortnnity. This proposed schedule Tvill more the planned launch date for Apollo 15 

to July or August 1971, maintaining our approximate six-month interval between launches on which the Apollo operations are now based. 

Review Board Report. 

1. Oxygen tank modifications 

probe during assembly 

and light alarm for single furl cell reactant valve closure 

valve closure 

priate, the hydrogen caution and warning trip lerel. 
58 

Our best current estimate, which may change as we proceed with the actions 

outlined abore, is that the modifications to the dpollo hardware which n-e now have identified will cost in the range of $10 to $15 million. 

I request approval of these actions at the earliest possible time. 

Rocco A. PETROSE, 

Bp07lo Program Director. 

DALE D. NYERS, 

Concurrence : 

dssociate dtltiiinistrator for Vaiinerl Space Flight. 

THOMAS 0. PAISE, Administrator. 

Approved : 

The C'HAIRXAX. Dr. Myers. 

STATEMEST OF DR. MYERS 

Dr. MYERS. bfr. Chairman, members of the committee, I appreciate 

the opportunity to present to you the response of the Office of Manned Space Flight to the Cortright Board report. I beliere the Board's 

recom~neiiclations are soulid, and you will find that we concur with 

each recommendation and are responding to each. Dr. Paine has discussed recoiiiineiidations 6, 7. aiid 8, which deal with SAS-i-wide implementation, for both manned and unmannecl activities. Dr. Petroiie 

vi-ill cover recommendations 1 through 6, which deal specifically with 

the Apollo program. I would like to speak to recommeiidation 9, which, 

for our other manned programs, including Skylab and our future program such as the shuttle and space station. 

Let me first read the recoinmelid a t' 1011: 

although directed towarcl Apollo, I beliere has important iniplic a t' 1011s 

The Manned Spacecraft Center should reassess all Apollo spacecraft iubsystenis, and the engineering organizations responsible for them at JISC and at its 

prime contractors, to insure adequate understanding and control of the en- gineering and manufacturing details of these subsystems at the subcontractor 

and vendor level. Where necessary, organizational elements should be strengthened and indepth reviews conducted on selected subsystems u-ith emphasis on soundness of design, quality of manufacturing, adequacy of test, and operational 

experience. 

First, we will expand the review to cover the three manned spaceflight centers-Manned Spacecraft Center, Kennedy Space Center, and 

Marshall Space Flight Center, for all Apo110 subsystems, including 

the command module, the lunar module, the launch vehicle and the 

ground support equipment. 

Second, n-e dl expand the review to iiiclude Skylab, and will use 

the results of these two studies and actions to set requirements for 

future programs. 

The reasessment of subsvstems vill be managed by the headquarters 

Office of Manned Space Flight and will include participation of Center 

top management aiid technical personnel most familiar with the subsystems. Additional personnel familiar with quality of manufacturing, 

engineering design aiid control systems, test and operations, and subcontractor management will be inr-olved., 

Although the history of Apollo missloll performance has been outstanding, we will reriew for each subsystem significant failure history 

trends, evaluate hazards to hardware in manufacturing, checkout or in 

flight, aiid will reevaluate possible "sympathetic failures" n-here reduiidaiit systems may lose their redundancy due to the failure of one 
59 

element affecting the perforinance of the second element, as occurred 

in the case of Apollo 13. Of course. in such a study we recognize fully 

that pressure vessels must not fail, because of their serious impact. 

Reviews will result in a report aiid presentation on each subsystem, 

n screening process, a further review indepth on selected subsystems, 

and a filial preseiitatioii given to an Office of Maiined Space Flight 

here in Washington. 

Throughout the review aiid presentation process, we will evaluate 

the adequacy of our engineering team at the centers, nt the prime 

contractors, and at the subcontractor level. 

A report on actioiis taken and those remaining on recommendation 9, 

as well as our actions on recommendations 6 and 7 will be presented 

to the Administrator by August 25. 

I believe that such a review will be extremely valuable in refining 

and developing a new level of underst:inding for Apollo and for our 

future programs. 

I would now like to call on 1)r. Petrone, Director of the Apollo 

program for the Office of Maniied Space Flight, to describe the 

actions to be taken by him relative to the recoinnieiidations 1 through 6, 

which deal specifically with the ,4po11o prograin. 

ST.\TEJIEST BY DR. PETROSE 

Dr. PETROSE. Thank 5’011, Jlr. Chairman. 

May I hare chart 1 (fig. el), please? This shows recommendation 

KO. 1 of the Apollo 13 Review Board. It addresses itself to removing 

froin contact with oxygen all virinp aiid the unsealed motors which 

can potentially short circiiit, and ignite adjacent materials or otherwice insure against a catastrophic electrically iiicluced fire in the 

tank. 

APOLLO 13 REVIEW BOARD 

RECOMMENDATION 1 

oTHE CRYOGENIC OXYGEN STORAGE SYSTEM IR THE SERVICE MODULE 

SHOULD BE MODIFIED TO: 

(a) REMOVE FROM CONTACT NTH THE OXYGEN ALL lv\i~~~~~ARD 

THE UVSEALED PvlQTOKS, NHICtl CAN POTENTIALLY SHORT 

CIRCUIT AND IGNITE ADJACENT MATERIALS; OR OTHERdISE 

INSURE AGAINST A CATASTROPHIC ELECTRICALLY INDUCED 

FI RE IN THE TANK. 

(b) MINIMIZE THE USE OF TEFLON, ALUMINUM, AND OTHER 

POTENTIALLY COMBUSTIBLE MATERIALS IN THE PRESENCE 

OF THE OXYGEN AND POTENTIAL IGNITION SOURCES. 

CHART #1 

FIGURE 21 
60 

The secoiid part of that recoinineiidatioii was to minimize the me 

of Teflon, aliiiniiiuin, and other potentially combustible materials in 

the presence of the oxygen and notentis1 irgiition colirceq. 

Cliart 2 (fig. 22), please. On this chart we sliow our redesign approach to the oxygen system. Onr objective is first to minimize flammable materials and ignition sources, and second, to retain our operational capability. 

Sow, specifically the inoclificationq we will make are in the heater 

assembl y-quaatity probe. We Fill eliminate the misealed fan motors. JT’e will replace the Teflon-coated wires with steel-sheathed 

wires. We will cliange the quantity gage from aluminum to stainless 

steel. We will eliminate heater thermal s\~itches. To give us an understanding, lioIyever, of the temperature on the heater elements, we will 

add a temperature sensor on that heater eleineiit. To improve reduuclaiicy in tlie tank once we take out the uwealed fan motor we will add 

a tliircl heater element. We also will modify the tank cap to simplify 

tlie basic assembly. 

We must also add n third oxygen tank to the service module to 

allow us to operate in the low-density regime, once we remove the 

unqealecl fans. 

This design was already in Ayork as a requirement for the Apollo 

16 axid subsequent missions. Kow, we can take that third tank, modify 

it as noted above, aiid introduce that into the service module. This 

will meet oiir requirements for Apollo 14 aiid 15 in the low-density regime. However, for Apollo 16 aiid subsequent missions where we hare 

REDESIGN APPROACH - OXYGEN SYSTEM 

0 OBJECTIVE 

MINIMIZE FLAMMABLE MATERIALS AND IGNITION SOURCES 

RETA I N 0 PE RAT I ONAL C APAB I LI TY 

0 SPECIFIC 

HEATER ASSEhlBLYIQUANTITY PROBE 

ELIMINATE FAN MOTORS * REPLACE TEFLON-COATED ”VRES NlTH STEEL SHEATHED “VRES 

CHANGE QUANTITY GAUGE FROM ALUMINUM TO STAINLESS STEEL 

ELIMINATE HEATER THERMAL SdITCHES 

ADD TEMPERATURE SENSOR ON HEATER SURFACE 

ADD TH I RD HEATER ELEMENT 

MODIFY TANK CAP TO SIMPLIFY ASSEMBLY 

ADD THIRD OXYGEN TANK 

0 ___- OXYGEN DISTRIBUTION SYSTEM 

ISOLATE TEFLON-COATED WIRES IN SHUT OFF VALVE 

CHART #2 

FIGURE 22 
61 

liigller oxygen requirements, we will have to review our oxygen 

subsystem to insure we can meet those requirements. 

Mr. GEHRIG. Dr. Petrone, will you eliminate the fans altogether 

or will yon just take them out of the tank’? 

I>r. PETRONE. In the module we are discnssing for Apollo 14 and 15 

we will eliminate them entirely. 

Mr. GEHRIG. There will be no fans? 

Dr. PETRONE. None on A4pollo 14 and 15. 

Atr. GEHRIG. HOW will you keep the-how will you pet ricl of the 

bubble in the tank formed by the heater? 

Dr. PETRONE. We have operated down to the regime of 35 percent 

and hare operated for periods of approximately 50 hours without 

the fan. This is R specific regime we hare experienced. But since we 

have not operated below 35 percent and since \re could have the 

difficulty you are referring to, we put in a third tank to keep US out 

of the regime in \I-hich we do not have experience. Howerer, on this 

flight we would expect to get that experience which requires zero “g” 

for a prolonged period of time. The only way we are going to get 

that experience is through flight testing. But I do note that for Apollo 

16, where our demands are even higher than Apollo 14 and 15, we will 

have to review the need for circulation and it may he necessary to 

introduce some other elements. 

However, we will not use unsealed fan motors. 

Senator GOI,DWATER. Dr. Petrone, I notice on your nest chart you 

call for unsealed fan motors, yet you do not show them. 

Dr. PETRONE. Those are to be reiiiored. I hare those as notes to 

myself. The basic chart would eliminafe the unsealed fan motor. 

In the oxygen distribution system, we will isolate the Teflon-coated 

wires in the shutoff valve. Here we found Teflon also in the high 

oxygen pressure area. 

On chart 3 (fig. 23), I show a basic drawing of what tlie tank 

would look like scheinatically with our modifications. The items on 

the left are notes which would say, one, we are going to reinore the 

unsealed fan motors. We will use sheathed wires~ The quantity probe, 

which we have to retain will be changed from aluminum to stainless 

steel. The thermal switches will have been removed. The temperature 

sensor, the second one will have been aclcled. As I mentioned, we will 

add a third heater element for redmidancy. The tank cap will be 

modified to simplify tlie basic installation of the heater probe 

assembly. 

Senator GOLDWATER. How can you keep this oxygen in a supercritical 

stage without some circulation ? 

Dr. PETROSE. Sir, we \i-ill clo that hy the heater. We will keep our 

pressure up to the 900-pomicls-per-scinnre-inch regime ancl in the high 

density areas, high cleiisity being the regime clo\rn to approximately 

35 percent, we feel we will get enough heat transfer dne to conduction. 

There 71-ill be no convection because yon are in zero “g.” and 11-e feel 

that Trill give US proper distribution. It is in a regime below 3.5 percent 

that we hare concern, that we has7e 110 experience, and for tlint i*cason 

we :ire :tdding a third tank for ,%pollo 14. 
62 

MODIFIED OXYGEN STORAGE SYSTEM 

0 UNSEALED FAN MOTORS 

SHEATHED WIRE 

0 QUANTITY PROBE 

.THERMAL SWITCH 

.TEMPERATURE SENSOR 

OHEATER ELEMENT 

0 TANK CAP 

FIQURE 23 

Senator GOLDWATER. What happens in minus (‘g.’’ conditions with 

the supercritical oxygen? It used to be circulated by fans but now you 

are depending on heat and you are not going to get it-you do not 

mint too much heat in it. How are you going to keep it moving? 

Dr. PETROXE. Sir, the concern comes in zero “g.”, in other voids, the 

forces, negative or positive ‘(g.” forces, are in your interest. They will 

help circulation. However, for zero gravity, the heat transfer due to 

conduction at the higher density regime will give us a distribution of 

heat and n-e can remove the fan as long as we are willing to add a 

third tank, to avoid the low-density regime. This is an area of the 

unknown, There are studies which would indicate to us we could go 

without that. However, we do not feel me want to take just theoretical 

studies. We want to see the performance of that at least in the 

Apollo 14 flight. 
63 

APOLLO 13 REVIEW BOARD 

RECOMMENDATION 2 

.THE MODIFIED CRYOGENIC OXYGEN STORAGE SYSTEM SHOULD BE 

SUBJECTED TO A RIGOROUS REQUALIFICATION PROGRAM, INCLUDING 

CAREFUL ATTENTION TO POTENTIAL OPERATIONAL PROBLEMS. 

RESPONSE 

@REQUALIFICATION NlLL BE ACCOMPLISHED TO: 

'VERIFY ADEQUACY OF DESIGN CHANGES 

' VERIFY REDUNDANCY 

DEMONSTRATE OPERATIONAL MODES 

CHART #4 

Chart 4 (fig. 24), please. This refers to recommendation 2 of the 

Apollo 13 Review Board, that the modified tank should be subjected 

to a rigorous requalification program. I might state that in the Apollo 

program we have a specific set of ground rules that me must meet on 

the ground, in our requalification. We dl requalify to verify adequacy of design changes, verify we do have adequate redundancy and 

demonstrate operational modes, working over many parts of the 

regime. 
64 

APOLLO 13 REVIEW BOARD 

RECOMMENDATION 3a 

.THE vZiARNlNG SYSTEMS ON BOARD THE APOLLO SPACECRAFT AND IN THE MISSION 

CONTROL CENTER SHOULD BE CAREFULLY REVlEvVED AND MODIFIED vVHERE 

APPROPRIATE, dlTH SPECIFIC ATTENTION TO THE FOLLOvVING: 

INCREASING THE DIFFERENTIAL BETJEEN MASTER ALARM TRIP LEVELS 

AND EXPECTED NORMAL OPERATING RANGES TO AVOID UNNECESSARY 

ALARMS 

RESPONSE 

@REVIE$/ INDICATES THATTHE HYDROGEN TANK PRESSURE vVARNlNG LIMIT 

IS THE ONLY ONE OF CONCERN 

*APPROVED MODIFICATION KIT NlLL BE INSTALLED $/HERE APPROPRIATE 

CHART #5 

FIDUEE 25 

Chart 5 (fig. 25), please. Here we come to a recommendation of 

the Review Board that has four sections to it. I will have a chart for 

each section. It basically addressed itself to the warming systems on 

board the Apollo spacecraft and in the Mission Control Center. It said 

these systems should be carefully reviewed and modified where 

appropriate. 

The first one addressed itself to looking at the differential that 

existed between master alarm trip levels and expected normal operating ranges. We have done that. 

Our review indicates that the hydrogen tank pressure warning limit 

is the only one of concern. There we have an approved modification kit 

to be installed in spacecraft where appropriate, to give us this wider 

range. 
A5 

APOLLO 13 REVIEW BOARD 

RECOMMENDATION 3b 

.THE UARNING SYSTEMS ON BOARD THE APOLLO SPACECRAFT AND IN THE MISSION 

CONTROL CENTER SHOULD BE CAREFULLY REVIEdED AND MODIFIED vVHERE 

APPROPRIATE, WlTH SPECIFIC ATTENTION TO THE FOLLOWING: 

*CHANGING THE CAUTION AND UARNING SYSTEM LOGIC TO PREVENT 

AN OUT-OF-LIMITS ALARM FROM BLOCKING ANOTHER ALARM WHEN 

A SECOND QUANTITY IN THE SAME SUBSYSTEM GOES OUT OF LIMITS 

RESPONSE 

0 BLOCKING OF CRITICAL SYSTEM PARAMETERS MINIMIZED BY PROVIDING 

SECOND LEVEL SENSING ON GROUND 

CHART #6 

FIGURE 26 

Chart 6 (fig. 26), please. Recommendation 3-B, here, referred to 

changing the caution and warning system logic to prevent an out-oflimits alarm from blocking another alarm when a second quantity in 

the same subsystem goes out of limits. 

Here we have had to resort to providing the second-level sensing on 

the ground in the Mission Control Center at Houston to minimize the 

effect of blocking of the critical system parameters. 

APOLLO 13 REVIEW BOARD 

RECOMMENDATION 3c 

0THE ,VARNING SYSTEMS ON BOARD THE APOLLO SPACECRAFT AND IN THE MISSION 

CONTROL CENTER SHOULD BE CAREFULLY REVIEWED AND MODIFIED NHERE 

APPROPRIATE, NlTH SPECIFIC ATTENTION TO THE FOLLOWING: 

*ESTABLISHING A SECOND LEVEL OF LIMIT SENSING IN MISSION CONTROL 

ON CRITICAL QUANTITIES NlTH A VISUAL OR AUDIBLE ALARM iiVHlCH 

CANNOT BE EASILY OVERLOOKED 

RESPONSE 

THE CAPABILITY FOR SECOND LEVEL SENSING IS BEING IMPLEMENTED 

0 A MASTER ALARM LIGHT IN ADDITION TOTHE PRESENT INDIVIDUAL WARNING 

LIGHTS IS BEING ADDED 

CHART #7 
66 

Chart 7 (fig. 27), please. Recommendation 3-C. This specifically 

talked of establishing a second-level sensing on the ground on critical 

quantities with a visual or audible alarm which could not be easily 

overlooked. That capability for second-level sensing is being implemented and we mill have a master alarm light in addition to the present individual warning lights. 

APOLLO 13 REVIEW BOARD 

RECOMMENDATION 3d 

0 THE WARNING SYSTEMS ON BOARD THE APOLLO SPACECRAFT AND IN 

THE MISSION CONTROL CENTER SHOULD BE CAREFULLY R E V I E d E D A N D 

MODIFIED vVHERE APPROPRIATE, WITH SPECIFIC AilENTION TO THE 

FOLLOViI NG: 

.PROVIDING INDEPENDENT TALKBACK INDICATORS FOR EACH OF 

THE SIX FUEL CELL REACTANT VALVES PLUS A MASTER ALARM 

inlHEN ANY VALVE CLOSES 

RESPONSE 

oMODIFICATION iivlLL PROVIDE A iNARNlNG IN THE COMMAND MODULE 

dHEN ANY ONE OF THE SIX FUEL CELL REACTANT VALVES CLOSES 

CHART #8 

FIQUBE 28 

Chart 8 (fig. 28). Recommendation 3-D. This addressed itself to 

providing independent talkback indicators for each of the six fuel cell 

reactant valves plus a master alarm when any valve closes. We will 

make a modification and provide a ma.rning in the command module 

when any one of the six fuel cell reactant valves closes. 

. 
67 

APOLLO 13 REVIEW BOARD 

RECOMMENDATION 4 

0 CONSUMABLES AND EMERGENCY EQUl PMENT IN THE LM AND THE CM SHOULD 

BE REVlEiNED TO DETERMINE A‘HETHER STEPS SHOULD BE TAKEN TO ENHANCE 

THEIR POTENTIAL FOR USE IN A “LIFEBOAT” MODE 

RESPONSE 

MANNED SPACECRAFT CENIER CONDUCTING EXTENSIVE REVIEA 

OF LIFEBOAT CAPABILITIES 

CHART #9 

FIQURE 29 

Chart 9 (fig. 29). Recommendation 4. This addressed itself to reviewing the consummables and emergency equipment in the lunar 

module and command module to enhance their potential for use in 

a “lifeboat” mode, This review is underway. We have a very extensive 

review going on and we will be getting reports within the coming 

weeks on what we can do in addition to what me mere able to do on 

Apollo 13. 

APOLLO 13 REVIEW BOARD 

RECOMMENDATION 5 

0 THE MANNED SPACECRAFT CENTER SHOULD COMPLETE THE SPECIAL TESTS AND 

ANALYSES NOYY UNDERWAY lh ORDER TO UNDERSTAND 1110QE COh’TJLkTELY THE 

DETAILS OF THE APOLLO 13 ACCIDENT. IN ADD!TIOh, Tt{€ LUNAR llODULE 

POiYER SYSTEM ANOMALIES SHOULD RECEIVE CAREFUL ATTPITlOh. OTHtR NASA 

CENTERS SHOULD CONTINUE THEIR SUPPORT TO hlSC IN THE AREAS OF AlvALYSlS 

AND TEST 

RESPONSE 

SPECIAL TESTS AND ANALYSES INITlATED BY INVESTIGATION iVlLL BE COhlPETED 

0 LUNAR MODULE POWER SYSTEM ANOMALY 

MOST PROBABLE CAUSE - LEAKING ELECTROLYTE RESULTED IN SHORT 

CIRCUIT 

SHORT C I RCU ITS 

SEALING hlODIFICATION >;IILL PRECLUDE FREE ELEClROLYTE FORMING 

OTHER NASA CENTERS ARE CONTINUING SUPPORT 

CHART #10 

FIQURE 30 
68 

Chart 10 (fig. 30) recommendation 5. This addressed itself to completing the special tests and the analyses now underway. Our response 

there IS those special tests and analyses initiated by the investigation 

will be completed. 

The second part of the recommendation addressed itself to the lunar 

module power system anomaly, saying that it should receive careful 

attention. The limar module batteries which JTere porn-ering. Apollo 13 

on the way home, did indicate a short circuit at approximately 97 

hours into the mission, for a rery short period of time 1 to 2 seconds. 

We have done a detailed analysis of the telemetry data that we had for 

that period of time and ve hare identified the most probable cause being leaking electrolyte which resulted in a short circuit. TT’e plan to fix 

thls by a sealing modification which will preclude free electrolyte from 

forming short circuits. 

The last part of that recommendation addressed itself to the other 

NASA centers which have given a very high level of support to the 

Manned Spacecraft Center in this area, to continue this support. Those 

centers will continue their support. 

Mr. GEHRIG. Dr. Petrone, the fix on the lunar module anomaly which 

is reported in the Board’s finding No. 47, this will not have any significant cost or schedule impact on the program, then? 

Dr. PETROXE. We do not see this as a large fix. It is a matter of adding 

potting compound to areas that vere exposed to possible electrolyte 

shorts and thereby sealing it. That mould not be a high cost fix. 

APOLLO 13 REVIEW BOARD 

RECOMMENDATION 6 

OWHENEVER SIGNIFICANT ANOMALIES OCCUR 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, FURTHERMORE, CRITICAL DECISIONS INVOLVING THE FLIGHTNORTHINESS OF SUBSYSTEMS SHOULD REQUIRE THE PRESENCE AND FULL 

PARTICIPATION OF AN EXPERT dHO IS INTIMATELY FAMILIAR WITH THE DETAILS 

OF THAT SUBSYSTEM 

RESPONSE 

PRESENT PROCEDURE 

*IN RESOLVING ANY ANOMALY RELIABILITY AND QUALITY ASSURANCE 

PERSONNEL REVlElN HARDNARE HISTORY FOR PRIOR ANOMALIES 

@AUGMENTATION OF PRESENT PROCEDURES 

0HARDlNARE HISTORIES iiVlLL ALSO BE REVlElNED BY TEST ENGINEERS 

OFINDINGS +/ILL BE PRESENTED TO PROJECT ENGINEER 

*START PRESENT SYSTEM FOR RESOLVING IN-FLIGHT PROBLEMS EARLIER 

.EXPERTS RELOCATED AS REQUIRED 

CHART #11 

PIQUBE 31 
t% 

Now, we will go to Chart 11 (fig. 31) recommendation 6. This recommendation addressed itself to insuring that whenever significant 

anomalies occur in a critical subsystem during final preparation for 

launch, standard procedures should require the presentation of all prior 

anomalies on that particular piece of equipment, including those which 

have been previously corrected or explained. Furthermore, critical decisions involving the flight ITorthiness of subsystems should require 

the presence and full participation of an expert who is intimately familiar with the details of that subsystem. 

In our present procedure, in resolving any anomaly, reliability and 

quality assurance engineers review the hardware history for prior 

anomalies. We are going to augment our present procedures by requiring in addition to the reliability quality assurances personnel, that the 

test engineer also review the hardn-are history and make a presentation 

of his findings to the project engineer. We presently have a system for 

resolving in-flight anomalies which starts with the final countdown. 

This is a grouping of cross-expertise in support of the Mission Control 

Center at Houston. We are now going to activate that system approximately 6 veeks before launch ~vhich will include all of the final launch 

preparations. This will bring together the cross-expertise that is needed 

in solving any one problem. 

Senator CERTIS. May I ask a question ? 

The CHAIRMAN. Senator Curtis. 

SUMMARY OF DESIGN PHILOSOPHY FOR FIRE PREVENTION 

Senator C~TIS. There is something I would like to have somebody 

comment on. When the fire occurred 3 years ago, the 204 fire, extensive 

modifications of the Bpollo hardware were made after that fire. Yet 

the A4pollo 13 Review Board found “A hazardous combination of materials and potential ignition sources in the oxygen tanks.” 

I would like to have you comment on that particularly as to any 

connection with the earlier studies of the fire. 

Dr. PETRONE. Sir,I would state that the review after the 204 fire 

concentrated on the cabin of the command module and the lunar module. The revielv came up TTith specifications that in effect, had to assume 

the fire would start or a spark would start, that it would be impossible 

to totally rule it out in a thing as large as the cabin, and then come up 

with materials specifications that would extinpish the fire in the 5 

pound per square inch pure oxygen that we have in space. 

The basic ground rule was to assume the fact that a fire would start, 

and determine how to contain it. Through the use of Teflon, which was 

the best material available, any fire that started in 5 pounds per square 

inch zero “G,” would smolder out, 

In addition, specifications and rules were arrived at on having firebreaks so that an ignition started in one place and maybe smoldering 

as it was going out would not set off other material nearby by having 

prescribed amounts that could be located in specific places. 

In this tank we have a different condition now. Instead of low pressure, we have this much higher pressure of a thousand pounds per 

square inch. In the design, the assessment was made that once we have 

the harness in and the cable in the sealed container, and have checked 

it out at a high voltage and see no short, then we will have no ignition. 

So that was the error, the abuse unknown at that time of overheating 
70 

this wire, as we did in Ithe improvised detanking procedure we did set 

up the potelltial then, of these arcs. Once you obtain ignition with the 

Teflon at this high pressure, there mas no containing it as there vould 

be in a cabin. 

Mr. GEHRIG. Dr. Petrone, I would like to follow that up with a question. After the Apollo 204 accident why were not all these subsystems 

revielyed, particularly where they involved an oxygen environment 

and where they had not been subjected to prior formal design review 

by NASA or Korth American Rockwell? 

Dr. PETROSE. I believe I can say the tank was reviewed. It Tvas 

reviejyed in the summer of 1967 by a group that NASA put in residence 

\Tit11 North American to review all systems and I believe the basic error 

vas made within the final criteria which assumed you could have an 

ignition .source and then in the cabin it would snuff out. In the tank, 

the testing of the liarness at the higher voltage was taken as proof that 

you would not have ignition. 

Rlr. GEHRIG. As I understand it, you did not test to see if Teflon 

would burn in a high density, high pressure oxygen environment. 

Dr. PETRONE. Well, there were tests made I believe at 20 pounds per 

square inch and thereMr. GEHRIG. Did it burn at 20 pounds per square inch? 

Dr. PETROXE. At 20 pounds per square inch it would smolder. It 

would not go out. hid the conclusion was that at higher pressures, the 

rate of combustion JTould not vary greatly. So, the knowledge that 

Teflon would burn in the higher pressures was acknodedged and 

is in our basic design specifications. The denial of an ignition source, 

that there would not be a way of setting it off in that tank, is the key 

failure. 

Mr. GEHRIG. Well, what bothers me about that is that that is precisely what the initial design concept was when they designed the 

cornmiand module. They said you have to have three things for a fire, 

an ignition source, oxygen and fuel and the way they were going to 

prevent fires in that pre oxygen environment was to prerent ignition 

sources. They did not prevent ignition sources and I thought that had 

taught everybody a lesson, yet here ve have another case where ve 

have a tank with a pure oxygen environment, we have fuel in the tank, 

and we do have ignition sources. 

I still do not understand \Thy NASA or the contractors did not test 

the Teflon under this high pressure oxygen and test for igliitioll 

sources. as 1 understand it from the Board’s report, the way that 

most of the testing is done is by striking and here the ignition source, 

and the ignition source was thought to be in the Spollo 204 accident, 

an electric arc which I understand concentrates the heat energy. Could 

you comment on that, please ? 

Dr. PETRONE. Yes. The testing is done on that harness in a sealed 

container and as such, understood not to get further wear or abuse by 

moremeiit as it would in a cabin. The test after it was assembled was 

collsidered to gire us assurance we would not hare an igllitioli source. 

?Jon-, through the abuse of overheating the wire we, in effect, set up 

an ignition source, and that is the error. 

Mr. GEHRIG. But this then is the same basic error that was made prior 

to the Apollo 204 fire. 

Dr. PETRONE. I do not beIieve I would consider it to be the same basic 

error. 
71 

Mr. GEHRIQ. Well, the same philosophy, design philosophy, was 

followed in the design of tlie command module prior to the 204 fire. 

The way to eliminate fires in a pure oxygen environment, as the testimony before the committee shows, as to eliminate ignition sources. 

,4s a matter of fact, they got pretty careless about what kind of 

material they put in tlie command module because they were so sure 

they did not have ignition sources. 

Dr, PETRONE. I referred here to the design in the tank, tested at a 

high enough 17oltage, revealed no ignition source. Kox, when the equipment is abused there is no ~~ay you can design it to know what it can be 

abused for. The unknown abuse set off a series of conditions which gave 

us the fire. 

(Additional information submitted for the record follows :I 

The following is a summary of the design philosophy adopted to prevent fire 

in Apollo following the Apollo 204 accident. 

FIRE PREVENTION IN APOLLO 

Following the Apollo 204 fire in January 1967, possible causes for spacecraft 

fires, and means of fire prevention were re-esamined. 

Three things must be present to start and sustain a fire: an oxidizer, a fuel 

and an ignition source. 

In manned spacecraft, an oxidizer (oxygen) must always be present to sustain 

life. Therefore, one look toward minimizing possible fuel sources, or toward 

eliminating ignition sources, in order to prevent a fire. Two different approaches were used to minimize the fire hazard in the cabin and in the oxygen tank, because 

the environment in these two locations is entirely different. 

Fire Prmention in the Cabin: The detailed examination of the Apollo space- craft cabin after the 204 fire soon revealed that it would not be possible to assure the elimination of all ignition sources. There are miles of wiring in the cabin ; there are hundreds of switches and connectors. And the potential for damaging these during checkout, or even during flight, always exists. A damaged piece of wire, a scraped piece of insulation, or a damaged connector can become a source of 

ignition. Therefore, the approach adopted for design and testing had to consist of 

eliminating as much combustible material as possible consistent with operation requirements ; protecting remaining combustibles against fires ; and arranging 

combustibles so that any accidental fire mould be self-limiting and of limited 

magnitude in the test, launch and flight atmospheres. In a low pressure oxygen atmosphere, many non-metallic materials become potential fuel sources ; however, metals will not burn. But some non-metals can also be found that will not 

bum at the 5 psi cabin atmosphere. Only Teflon wire insulation was permitted. Space suit covers were changed from Nyon to Fiberglass cloth (Beta-cloth) . New 

potting compounds that do not burn viere invented. All nylon and polyurethane plastics were ruled out, and were generally replaced with fluorine-based com- pounds that will not burn in this atmosphere. Then tests were made. In these tests, prorocative ignition was deliberately 

attempted all over the spacecraft. Snd it was demonstrated that in the 5 psi cabin atmosphere the spacecraft fire could not be sustained; it always selfextinguished. 

It was also found that the same materials that were safe in the 5 psi (space) cabin atmosphere would burn at 16 psi in pure oxygen-the atmosphere that 

had been planned for the launching pad. We therefore clhanged from a pure 

oxygen atmosphere to one that contained 60% oxygen and 40% nitrogen on the 

launching pad. Tests again demonstrated that fire would not propagate in the 

newly rebuilt spacecraft in this atmosphere, even at 16 psi. 

The oxugen storage tank: Inside the oxygen storage tank, the pressure is 

900 psi. The materials that were found to be safe in the cabin would still burn 

in this atmosphere. Even metals. aluminum. steel and inconel. can be ignited. Therefore, it was not possible to remove all the “fuel”; instead, the ignition 

Source would have to be eliminated. 

Inside the tank, the situation is far more restricted than in the cabin : there 

is only a small amount of wire, and a few electrical components. These could 
72 

conceivably become an ignition source through overheating, or through arcing. The possibility of overheating was eliminated by installing fuses in the circuit. 

This left only arcing as the potential igniton source. It was known that a damaged wire or damaged insulation would be needed 

to strike an arc. And It was reasoned that if it was demonstated, after the tank was assembled, that the wiring was not damaged, no damage could occur later 

on because the tank was sealed. The wiring was checked after assemblythrough the use of a dielectric strength test-and found to be undamaged. Of course, we now know that one important factor was overlooked in this 

reasoning : the miring inside the sealed tank could be subsequently damaged 

through exteral sources. This is just what happened during the special de-tanking procedure prior to launch. The wiring was damaged, a potential ignition source was created, and the Teflon insulation was ignited when the arc occurred 

in flight. 

For future flights, potential ignition sources for the materials as located within the tanks are being eliminated by encasing all mire in stainless steel and by 

removing the fan motors from the tank. 

Senator HOLLAND. Mr. ChairmanThe CHAIRMAN. Senator Holland. 

OXYGEN TANKS 

Senator HOLLAND. I have several questions. First. why was the second oxygen tank a casualty after the accident had happened in the 

first tank? 

Why mas that sequence apparently unavoidable ? 

Dr. PETRONE. I do not believe WQ fully have explained that. The 

analyses have considered failurs of the second tank. It took about 

an hour and a half to lose its pressure. I do not believe we fully understand exactly how the first one caused the second one to lose pressure. 

I wonder if Mr. Cortright might care to comment on that from the 

Board’s standpoint. 

Senator HOLLAND. It seems to me this is one of the critical points. 

You had two tanks in order to have one available if there was an 

accident on the first, and yet the accident on the first was communicated to the second so as to take it out of usefulness; and my question 

is why8 

Mr. CORTFUGHT. Senator, there are at least two possibilities here. 

There was a severe shock or jolt :tt the time the first tank which actually was tank No. 2, ruptured, This rapid expulsion of oxygen built 

up a pressure in the bay that exceeded 20 pounds per square inch 

and blew a rather stlrong structural panel off in n few thonqandths of a 

second. The resulting jolt mas sufficient that it might have jolted open 

the relief valve to oxygen tank No. 1 and that may have failed to 

reseat properly, in which event the tank would slowly leak down as 

it did. 

Another possibility is that the metal elements,) primarily the tubing 

in the cap to tihe top of the oxygen tank No. 2, could have been thrown 

violently aside when the top of the tank ruptured, struck a line to 

oxygen tank No. 1 and caused a small leak in that line. It would take 

only a small leak to deplete the oxygen in the time it mas depleted. 

Now, we found no way to prove by testing which of those two, if 

either, was the cause. Virtually all other known or observed events 

were demonstrated in subsequent tests. 

Senator HOLLAND. What additional precautions are you t a k’ ingnow 

to do away with any possibility of a similar communxation between 

tanks one and two or even to tank three, which you propose to add ? 
73 

Dr. MYERS. We are going to review all of the possible sympathetic 

damage conditions in the Apollo spacecraft and its launch vehicle, 

but it is very clear that the protection against damage from an exploding ta.nk or rupturing tank to another tank close by is one that 

we just cannot design against. We must keep from having tank ruptures. That is a fundamental premise of design that we must follow in 

these designs. So, we are going to review other systems to see whether 

there are other elements other than tank failures that could give US 

this interaction from one redundant system to another, but it is clear 

that in the tank area we just have to protect against tank failures. 

Senator HOLLAND. Why does the addition of EL third tank, as you 

now propose, obviate any possibility of just such an accident affecting 

all threa tanks? 

Dr. MYERS. Of course, we are here removing all Teflon coated 

wiring, any of the kinds of materials that could give us difficulty, from 

these tanks and we are designing now with every precaution that has 

come out of our Board’s report to be sure we do not have a tank 

rupture. 

Senator HOLLAND. Does the addition of a third tank bring on any 

space problems within the confines of the vehicle? 

Dr. PETRONE. No;it does not, Senator Holland. As I mentioned, 

ve had planned to put a third tank in Apollo 16. We have an area 

called bay 1 wherein with Apollo 16 and subsequent, we will add also 

:L third hydrogen tank and a scientific instrument module base. We 

will not have a problem on weight or space as far as adding a third 

tank for Apollo 14. 

WIRE COVERING 

Senator HOLLAND. If I understood your testimony correctly, you 

said you vould obviate the possibility of the fire hazard in certain 

parts of the wiring by using stainless steel covering for your wiring. 

But you still are continuing teflon covering for some of the wiring. 

Why the differencel 

Dr. PETRONE. I do not believe we have any Teflon-covered wires. 

They will all have been removed. The stainless steel sheathed vires 

will be used for any wires that have to go into that tank, 

Senator HOLLAND. I understood some earlier testimony to indicate 

that certain parts of the wiring would remain Teflon-covered. Was I 

mistaken ? 

Dr. MYERS. Senator, this wiring, for example, that goes into the 

tank and down inside the oxygen probe will all be removed, all that 

Teflon-corered wiring Tvill be removed and me will use only the stainless steel sl~ea~thed wire. 

Senator HOLLAXD. Where will the teflon wiring continue to be used? 

Dr. MYERS. There will be no Teflon wiring used in the tank at all. 

The remaining pieces of Teflon are insulators that insulate the capacitance gage from the grounding of the outside of the tank. These 

pieces of Teflon are inert and since these wires are all removed, there 

are no ignition sources any1-i-here near them. We are continuing to 

review the possibility of removing those Teflon pieces but since there 

are no-since all the electrical sources inside the tank now will be 

covered with stainless steel, we believe that through our testing we 

will be able to prove to ourselves that that piece of Teflon is satisfactory. 
74 

Senator HOLLAND. Well, why retain any Teflon wiring if you know 

Dr. MYERS. Not TeflonSenator HOLLAND (continuing). And that the stainless steel-covered 

wiring is not flammable? 

Dr. MYERS. I am sorry. We xi11 not have any Teflon-covered wiring 

inside the tank. There is a Teflon insulator for this capacitance gage, 

here and here. Those are the only remaining pieces of Teflon in the 

tank but not Teflon-covered viring. 

Senator HOLLAND. Well, it will be flammable wherever it is; will it 

not 1 

Dr. MYERS. Yes. The problem in this capacitance gaging system is 

that me have a balanced dielectric insulator betvyeen the ground of 

this tank and this inner capacitance gaging system, and although we 

are looking at possible other materials for that capacitance gage system we hare not found one that is satisfactory. Since we have removed 

all electrical wiring from the oxygen source, we believe ve can continue with this design of the capacitance gaging system. 

that that is flammableCOUNTDOWN PROCEDURE 

Senator HOLLAND. Rly next question is this : Will your countdown 

procedure include new tests applicable to the oxygen tanks which have 

not been used heretofore? 

Dr. PETRONE. I guess I would like to address that-you say new tests. 

There \Till be new procedures in terms of the new design that we are gojnp to have. With the modified design, we will bring it up to pressure 

differently. However, the use of the higher yoltage which was done to 

shorten the time required to pressurize, m-e still, and this is part of the 

review, may have to continue that, It is a matter of some 3 to 

4 hours of extra time that we would have to weigh. We do not 

count the higher voltage to have been the problem except on this improvised detanking procedure. That, of course, was a procedure that 

did cause the difficulty. It was not the tanking procedure. It was Then 

w0 ran into trouble detanking that an improvised procedure did 

allow the higher voltage to be used with the inadequate thermal srritch 

which did not open up when the limit of approximately 80’ was 

reached. 

Senator HOLLAND. Well, my question I will renew. Are yon going to 

have new procedures involved in the countdown which will guard 

against a recurrence of that kind ? 

Dr. PETRONE. Yes,we will. We will have new procedures based on 

the nem. design. We will. 

APOLLO FUNDING IN RnLATION TO NEW SCHEDULE 

Sena;tor HOLLAND, I have two practioal questions. The House reduced the fiscal year 1971 appropriation for research and development 

by $106.1 million below your request, indicating that the cut should 

be taken in tlhe Apollo program. In fact, the House Appropriations 

Committee in its report, recommends that the ,kpollo 14 flight be deferred until n&er the first of calendar gear 1971 and states : “The funds 

recommended \Till provide for one Apollo flighlt instead of kwo flights 

in fiscal ymr 1971.” 
75 

Your decision, now announced to this committee, not it0 launch 

hpollo 14 until at least January 31, 1971, which moves the planned 

launoh dalte for ,ipollo 15 to ,July or Angust 1971 at the earliest, agrees 

with the Honse recommendation. Will {this reduce the cost of the 

A4.pollo lprogr*am for fiscal year 1971 ? 

Dr. MYERS. Onr estimate is that change in the schedule would save 

approximately $20 million rather than the $106 million. 

Senator HOLLAND. Why the difference between the figure thak is 

stated by the House report and action and at the figure now stated by 

Dr. MYERS. I think the iproblem probably is in the degree of operations at the Clap that are involved and the degree of support from the 

contractors. Tt7e all are dealing with that problem now of ho\T to retain 

the capability of top svstems engineers in snplpo& of our programs at 

the Cape, system engineers within our centers, and the people that 

really understand these swtems at the contractors. When m-e look at 

delays in schedules it is difficult to estimate what kind of capablity we 

must maintain in support of these launch operations. 

Senator HOLLAND. Let us see if I understand what you have said. 

You are suggesting something different from the HOUS~, that instead 

of the reduction made by the House of $106.1 million, their ramn 

given that thev expecrt mly one Apdo flight instead of two flights in 

fiscal 1971, what redaction do you nov- recommend ? 

Dr. MYERS. We believe that you could save approximately $20 million on that mom of the schedule but as we testified here today, the 

changes to the tank itself, we believe, x-ill cost something of the order 

of $10 to $15 m5llion. So, it really is a kind of tradeoff as far ads the 

funding is concerned and, as we testified, we believe we can absorb 

these changes \Tithin lthe Gpollo budget for this year. 

Senator HOLLAKD. And the reason that you can absorb them is due 

in part to the fact that you are slowing down the ,4po11o launches, is 

that correct 1 

Dr. MYERS. Yes, sir. That does trade off, approximately even. 

Senator HOLLAND. These questions are going to be very troublesome questions in the appropriations process and perhaps particularly 

so in thk confereiice and as 1 naw understand it, yon think that about 

$20 million is the maximum saving that would result and that $10 to 

$15 million of that would necessarily be used in making these changes. 

Dr. MEYERS. Yes, sir. 

Senator HOLLAXD. To avert any possible recurrence of trouble vith 

the oxygen tank. 

Dr. MEYERS. Yes, sir. I think there is another point, Senator, that 

might be added here. In my testimony I discussed the further thorough 

inrestigation of our engineering teams and the capability of those 

teams to respond and to be on site properly and to understand in 

depth the kind of details that we really find we must understand in 

these sytems. hs the p pol lo program schedule is either delayed or 

stretched. the challenges that we have for top technical people is soinething we must continue to reinforce to be sure that xe have the proper 

support to these flights. And as a result of this review, one of the indications might be that we may vant to reinforce the capabilities of some 

of our subsystem activities to be sure that we are properly supporting 

these flights. If that were the case, it would perhaps make it even 

tighter for Apollo problems for this coming year. 

you ? 
76 

EFFECTIVENESS OF COUNTDOWN AT TWO LOCATIONS 

Senator HOLLAND. Now, my last question is one which comw to me 

from various of the personnel in the Kennedy Space Center complex. 

Is there any reduction in the effectiveness of the countdown due to the 

fact that the countdown takes place at the Kennedy Space Center 

whereas the Manned Space Center is over at Houston ? 

Dr. MEYEFK No, sir, I do not believe so. I have seen tremendous communications carried on between those two centers and from what I 

have seen of the balance of effort between the two centers, it has turned 

out to be a very fine reinforcing kid of an activity. The testers in the 

field questioning at times the people who technically mere involved 

in the development of the equipment has given us a balance of discussion that has been good as far as understanding thesesystems. 

Senator HOLLAND. Well, I think you can understand why there is 

still skepticism on this question present among some of the people sta. 

tioned at the Kennedy Space Center. They have never felt that removal so far apart of the Manned Spaceflight Center-the mission 

control center-at Houston from the launch operations center at Kennedy Space Center vas a wise thing and could possibly result in the 

degree of efficiency that vould have prevailed if the two centers had 

remained at the same location. 

Thank you, Mr. Chairman. 

The CHAIRNAN. Go ahead. Have you finished your statement Z 

Dr. PAINE. I believe our statement is concluded, Mr. Chairman. 

The CHAIRMAN. Senator Goldwater ? 

We will be happy to respond to any further questions. 

FIREPROOF MATERIALS 

Senator GOLDWATER. Yes. 

Dr. Petrone, on chart 10 (see fig. 30) you indicate that short circuits have been caused by free electrolyte forming. How could that 

happen in that atmosphere? 

Dr. PETRONE. These batteries have electrolyte in them as a process 

of charging. We also have to vent batteries to allow hydrogen gas 

which forms as part of the current forming process to escape. In this 

venting mechanism there is the potential for free electrolytes to escape and form in a globule. Despite the fact that we are in weightless 

condition at that time, some of the pressure that could build up could 

also expel some of the electrolytes, then letting it go into the terminal 

area where one then could get these short circuits. 

Senator GOLDWATER. Are these cadmium batteries ? 

Dr. PETRONE I believe they are similar. 

Senator GOLDWATER. The reason I am interested, I recall one day 

landing just behind another jet and that jet blew up because a cadmicm battery blew up. I believe it had been allowed to get dry and 

I imagine the same process took place. I was very interested to ask 

that question because I am acquainted with electrolysis but I never 

heard of this but it certainly could happen. 

Dr. Paine, we hear a lot about materials that will not burn and they 

are called fireproof and then they do burn and they are not fireproof; 

for example, Teflon. Now, I could haw told you a long time ago that 

Teflon burns because my 11-ife tells me every time I use a frying pan I 

burn it. What do you mean by “burning” and %reproof”? 
77 

Dr. PAINE. That is a very good comment, Senator Goldmater. The 

burning process, of course, is what you really fundamentally call in 

simplistic terms an oxidation. It is important to realize that every 

place in which a strong oxidizer is stored in any material, there is 

the possibility of tn-o chemicals being in contact and reacting with each 

other. 

In the case of Teflon the great resistance of this material to oxidation is perhaps best demonstrated by the fact that we did keep it for a 

substantial number of hours at Cape Kennedy during this detanking 

procedure at a heater temperature of 1,000" in contact with very high 

pressure oxygen without igniting it at that time. We degraded the 

Teflon badly and damaged it sufliciently so that we set up this later 

sparking ignition. Teflon belov about 1,200' will not ignite even at n 

thousand p.s.i. oxygen, although once ignited Teflon in any pressure 

above about 20 p.s.i. will continue to oxidize, to burn in pure oxygen, 

as was brought out in the questioning of hlr. Gehrig. 

What me have to do in the case of a spacecraft obviously, is to 

contain strong oxidizers and strong reducing agents within materials 

in such a way that we sufficiently minimize the chance of any reaction. 

But if the temperatures are raised high enough, of course, then they 

will ignite. Almost any time you get two dissimilar materials in 

contact they will begin to react against themselves. 

Senator GOLDWATER. Did you know this before the tank accident Z 

Dr. PAISE. Yes; as a result of the followup of 204 fire extensive 

tests were made of many materials. MTe found down at the five p.s.i. 

pressure at which we operate our cabins, there is a good deal of safety 

and with Teflon we can extinguish and block and have an opportunity 

to contain any ignition even if it begins. We found above 20 psi. 

we get the opposite effect. The Teflon then continues to burn. ,4nd the 

high p.s.i. as we have in the tank, Teflon and even metals like iron and 

steel and aluminum, will burn, once ignited. 

Senator GOLDWATER. I have heard that NASA is switching flight 

suits made of Nomex, whatever that is, to flight suits made of durette. 

Would you explain what Nomex is and dl it burn ? 

Dr. PAINE. Those particular terms are not familiar to me. 

Senator GOLDWATER. They are not in your dictionary either 1 

Dr. PAINE. Perhaps Dr. Low will comment on that. 

Dr. Low. Senator, Nomex is a ny.lon-type fabric that has a high 

temperature resistance property in air. It mill not burn, only char, 

while nylon will melt at a much lower temperature. 

Senator GOLDWATER. At about what temperature? 

Dr. Low. Nomex will char at 840' F. but nylon will melt at 482O F. 

Seantor GOLDWATER. Is this the same material that race drh-ers wear 

Dr. Low. Yes, sir. 

Senator GOLDWATER. Will it withstand gasoline fire temperatures? 

Dr. Low. No, sir. 

Senator GOLDWATER. It will for a short period of time. 

Dr. Low. Yes; for a few seconds but a better material has been 

dereloped, that has a higher flame-resistant temperature than Nomex. 

TTe are IIOW using the material for the astronauts' aircraft flight 

corcralls and im-estigating its use for firefighting uniforms. 

today ? 

Senator GOLDWATER. Is that what you call durettc? 

47-470 0-.70--6 
78 

Dr. Low. Yes. 

Senator GOLDWATER. We saw some demonstrations in this room of 

the material that you developed at NASA that prevents burning. None 

would burn under even the-I think the temperature was 3,000". 

Dr. Low. The best material TTe have in the space suits are made out 

of what is a beta cloth which is a fiberglass made in a special my so 

that the glass fibers themselres will not irritate the skin. However, 

that beta cloth fiber does not wear very well. It wears out quite quickly. 

The Teflon materials and the astronauts' space-flight coveralls made 

out of Teflon material are not quite as fireproof but will not burn 

ordinarily in the air and this Teflon material does have much better 

wear capability than the beta cloth. 

Senator GOLDWATER. Thank you very much. 

Mr. Cortright, in your testimony before the House you said : 

I think we ran into a new phenomena that was not widely recognized before 

and that is that Teflon can be ignited rather easily if an electric arc is the igniting 

mechanism, and we ran tests to show that these small amounts d electrical 

energy were sufficient if they were in the form of an electric arc which concen- trates the heat very locally in the material, and we ran additional tests to show 

that even through the one amp relatively quick low fuse that was on the line to 

protect it, you could get energy ten to 100 times in excess of what was required 

to ignite the wire insulation. 

Am I right in saying that the ignition temperature of Teflon is 1,300, 

about 1,300"? 

Mr. CORTRIGHT. Ithink that is about right, Senator Goldwater. 

Senator GOLDWATER. And you can get that temperature out of 1- 

amp arc? 

Mr. CORTRIGHT. Yes. sir, in a very local region you can get higher 

temperatures than that. The arc itself generates extremely high temperatures and it requires that the heat is so concentrated that it requires little total energy. 

Senator GOLDWATER. What is the voltage used on that particular 

system ? 

Mr. CORTRIGHT. The voltage used in the spacecraft is 28 volt D.C. 

Senator GOLDWATER. A11through it? All systems? 

Mr. CORTRIGHT. On the motors it is 115 a.c., pardon me. The basic 

power supply in the spacecraft is 28 volt D.C. Rut the fan motors were 

the higher A.C. voltage. 

Senator GOLDWATER. Are there any special conditions needed to get 

a concentration of energy in an arc or will it always occur! 

Mr. COKTRIGHT. Well, I am not sure how special they have to be. I 

can tell you the sort of arcs that we struck in order to get this to 

ignite. For example, if you skin the insulation on a wire to where it is 

just exposed to wire and carried a single strand of a multistrand vire 

into that skinned portion and then struck the arc by first melting that 

little piece of wire and then having the arc strike across the ionized 

gap, that would start it very well. 

On other occasions w0 created the short by means of a pointed screwdriver which robably or definitely would not melt and ionize as readily 

what you really need is a point contact in the presence of a very thin 

shaving of insulation such as you get in a skinned area. Snd that vas 

sufficient to start ignitions with less than 5 joules of energy. 

Senator GOLDWATER. I imagine the fuse blew out in that system. 

as the single i? lament of wire and that also started the fire. So, I think 
79 

Mr. CORTRIGHT. Yes. 

Senator GOLDWATER. It did ? 

Mr. CORTRIGHT. The fuses blew after the short. 

Senator GOLDWATER. HOW fast a fuse do you have? 

Mr. CORTRIGHT. TTell, it is a quick blow fuse but it n-as, I believe, at 

the S~OTT end of the quick blow range. I cannot give you the time but it 

s as sufficient to pass-excuse me. Dr. Clark remembers that it took 

about 30 milliseconds of 8 amps to blow and we knov that this 

could pass at least 20 joules of energy and perhaps as much as over 100. 

Senator G~LDWATER. Was there any vay that that could have been 

discovered during the Apollo 204 investigation ? 

Mr. CORTRIGHT. Had someone thought to run that particular type of 

ignition test, it would have been discovered, yes, sir, because that was 

not a highly complicated test to run. It was merely a matter of suspecting that this could occur. 

Senator GOLDWATER. In your statement you say that one.series of 

tests demonstrated electrical ignition of teflon insulation 111 supercritical oxygen under zero “G” and one “G,” and provided data on ignition energy and burning rates. 

Did you run similar tests 011 5 p.s.i. pure oxygen atmosphere under 

one and zero “G” conditions ? 

hlr. CORTRIGHT. Not as part of these investigations, but I believe 

those tests were run after the Apollo 204 fire. 

SUPERCRITICAL FLUID 

Senator GOLDWATER. Just one more question. Maybe two more questions. You talk about supercritical state of oxygen. For the record, 

could you-could someone define what you mean by supercritical ? 

JIr. CORTRIGHT. Supercritical fluid is a single phase fluid. Normally, 

for example, a container of vater, if I had a top on that, you would have 

liquid water and water vapor above it with a plane of separation. This 

is two-phase. Liquid oxygen in a container would look like that also but 

with liquid oxygen here and gaseous oxpg-en above it but in a certain 

r.egime of pressure-temperature combination that parting plane would 

disappear and the fluid would totally fill the tank in a single phase condition and it is maintained that way throughout the Apollo mission. 

Senator GOLDWATER. It would not then be really a fluid, would it ? 

Jlr. CORTRIGHT. Yes, it is a fluid. 

Senator GOLDWATER. With gasMr. CORTRIGHT. Well, Senator, you can have some very interesting 

discussions on whether it ils more liquid or more gas and we nerer 

quite did resolve that. It is more like a very, very heavy gps that 

is more dense than liquid or as dense as liquid or a liquid that behaves like a pas. You can take your choice. 

Dr. PA\l~r. A very stable fog \~ould be another way of putting it. 

Senator GOLDWATER. A very what ? 

Dr. !?.USE. h very stable fog that would completely fill the co11- 

Senator GOLDWATER. Would the term “cryogenics” be compatible 

Dr. PAINE. In order for oxygen to be supercritical it does require 

t a her. 

with supercritical? 

cryogenic temperatures, so that is a part of it. 
80 

Senator GOLDWATER. Dr. Petrone, on figure 28 you talk about providing independent talkback iiidioatiors for each of the six fuel cell 

reactant valvea. I thought all of the spmships had recorders that 

the command pilot could see or that could be reported to earth. 

Dr. PETRONE. In the particular oase of the fuel cell valves, the indicator system was so wired that both the oxygen and the hydrogen 

valve had to close before an indication was giren. This is what we 

have modified or are in the process of modifying now so that a warning will occur if either one or the other should close as happened here 

due to shock. There was in this particular flight not much that could 

be done because m-e had lost the oxygen supply, but in the future 

we would want to know that if either one of the oxygen or the hydrogen valves closes-he will get a light and a horn to Tarn him. 

Senator GOLDWATER. I am rather surprised that you did not have 

this before. One light for hydrogen would not be of much use in trying to tell which one you lost. 

Dr. PETROXE. It turns out the way it is TF-ired up, when you actuate 

a single switch you energize both valves. It is tied into the may you 

bring the system up. A single s\\ritch command will give both valves 

the command to open. That was the way that the indication came out, 

but we have changed that as a result of the review that \\re made after 

the Apollo 13. 

Senator GOLDWATER. Was it a single switch, a single double-throw 

switch for two instruments, was it just a single sn-itch for one instrument on two valves, it that aorrect ? 

Dr. PErRom. Yes, sir. Normally in flight you would never have to 

throw that valve. You would never want to bring the fuel cell down. 

You have lost it then. So, it was a type of ;L switch that on the groud 

was needed to bring the fuel cells up and get them operating properly, 

and in flight one would never want to close that. 

Dr. PAINE. An analogy that would be very familiar to you, Senator Goldwater, is that ve had the indications essentially as you have 

with both wheels up in an aircraft, and what happened was that essentially in this case we had one wheel down and ve did not get an 

iiidioatioii of which was which. We TF-ill now have an iiidicdion such 

that we can see either valve being opened. 

Senator GOLDWATER. Thank you. 

Thank you very much, Mr. Chairman. 

The CHAIRXAX. Senator Smith? 

Senator SMITH of Illinois. No questions. 

The CHAIRMAN. Senator Holland ? 

Sbnator HOLLAND. No further questions Mr. Chairman. 

The CHAIRJTAN. Mr. Gehrig. 

THERMOSTATIC SWITCH 

Mr. GEHRIG. Thank you, Mr. Ghairman. Mr. Cortright, you have 

said-in testimony before the House committee-that the thermostatic 

switch discrepancy vas not detected by NL4SLA, North American 

Rockn-ell, or Beech in their review of (documentation nor did tests 

identify the iiicompatabilty of the svitches with the ground support 

nor acceptance testing required switch cycling under load as should 

have been done. It was a serious oversight in which all parties shared. 

equipment at the Kennedy Space Center since neither qualific a t’ 1011 
When and how did the Board detect the thermostatic switch 

discrepancy 1 

Mr. CORTRIGHT. There Tvere tests being conducted to determine 

what damage might have occurred to the wiring which ran through 

the tank due to the extended heater operation. 

Now, this test was first run xitli the switches \Tired closed, in 

other words, wired out of the circuit, assuming that the switches 

might have failed closed, although we did not really hypothesize at 

the time how this might have happened, and this showed that the 

wiring was damaged as mentioned in the Board’s report. 

Subsequently, then, it was reasoned that, well, the history of the 

switches was such that they had a tendency to fail to open ; that mas 

the chronic problem witih them. Therefore, we decided to put the 

switches back in the circuit and determine whether or not the wires 

Twuld be damaged even if the switches were in for this extended 

operation. And it was at that point in time that the switches failed. 

In other words, it was not a test specifically to test the switches, but 

rather a test in which the switches xere used and failed; this was 

hov~ rve discovered the sxitch discrepancy. I think that answers your 

quest ion. 

Mr. GEHRIG. Yes. Do I understand that people did not realize that 

the switches were rated for 30-volt, direct current, and that there 

.irere 65 wlts on the line? 

Mr. CORTRIGHT. That is correct. 

Mr. GEHRIG. Is that correct? 

Mr. CORTFLIGHT. Yes. 

Mr. GEHRIG. Who designed the tank? 

Mr. CORTRIGHT. Beech. 

Mr. GEHRIG. Beech designed it. Now, Beech, as I understand it 

from your report or from your testimony before the House, uses 

65 volw, alternate current, at their plant to test. 

Mr. CORTRIGHT. That is correct. 

Mr. GEHRIG. Would not the engineers recognize that they had 28 

volt sn-itches and that might cause a problem? 

Mr. CORTRIGHT. Itturns out that the switches were quite capable 

of operating at 65 volts alternate current in the manner in xhich 

Bee& used it. 

Mr. GEHRIG. Isee. And they did not know that there were 65 volts 

direct current down at the cape? 

Mr. CORTRIGHT. Well, they were notified that there was in the specification witten by North American Rockwell. 

Mr. GEHRIG. But they did not connect this with the 30-volt switches 

actual1 , as I understand it, 

Mr. 8 ORTRIGHT. That seems to be the case. 

IGNITION SOURCES 

hfr. GEHRIG. Dr. Petrone, going back to the discussion we were 

haling before, there is an awful lot of oxygen on the Apollo command 

and service modules, on the lunar module, and on the Saturn V 

booster. Is NASA4 now going to take a look at all of these oxygen 

sources, look for ignition sources and combustible materials in the 

vicinity to make sure that they eliminate this hazard? 
82 

Dr. PETRONE. Yes, we are. We have those tests now underway and bhe 

thing we want to assure ourselves is that the materials that are there 

are tested under the conditions which they see. In other words, there 

are many types of tests that they call impact sensitivity tests, many 

different ways to qualify material to go into a lox tank. For example, 

we now, as a result of the Spollo 13 experience, review all our criteria 

to assure not only tihat we meet certain specifications but we meet all 

the conditions that exist in that tank. That activity is now unclern-ay. 

Mr. GEHRIG. And are these people going to look at ignition from 

electric arcs rather than just from impacts and things like that ? 

Dr. PETROKE. Yes, n-e are. 

Mr. GEHRIG. Because this is very similar, it seems to me-I looked 

back at the testimony before the commit,tee on the Apollo 204 accident 

ancl it reads very similar to what we have today. Testimony at that time 

was: "The approach to fire prevention is to prevent the ignition of 

combustion by attempting to remove a11 possible sources of ignition." 

That was the design philosophy before the 204 fire. 

FURTHER TESTS 

Mr. Cortright, on page 5-1 of the report of the Bpollo 13 Reviex 

Board, it is stated that, "Further tests and analysis which will be carried out under the overall direction of the Manned Spacecraft Center 

will continue to generate neTr information relatiJ-e to this accident. It 

is possible that this evidence may lead to conclusioiis differing in 

detail from those which can be drawn now." Are these tests and analyses complete and if so, has any evidence dereloped that might lead 

to different conclusions than those presented in the report ? 

Mr. CORTRIGHT. Mr. Gehrig, the tests and analyses are not yet 

complete but as yet have not led to any different conclusions. I can tell 

you briefly at least tnTo of them that are still underway. For example, 

at Beech Aircraft Corp. there is essentially a duplication of the detanking procedure on ;L flight quality tank which is going on for the second 

time. The first time the switches failed in a little different manner and 

hence, it is being repeated Kith the switches closed and actually the 

switches are out of the circuits, thereby simulating the manner in vhich 

is failed at the Cape. 

Currently it is not possible to insure that the switches will fail closed 

because one of them, for example, melted out the contact points and 

they dropped out and it failed open, although we know that the two at 

the Cape failed close. 

That is one test that is being completed this week. 

Another series of tests and analyses is to further refine our understanding of the manner in n-hich the panel ruptured and blew out. 

The first series of tests that were run were conducted with panels without cutouts in them and now we are attempting to account for the 

presence of cutouts in the panel which were present in the actual flight 

hardware and this may bring our experimental analyses and test data 

into better agreement with what we know happened in space. 

TANK TEMPERATURE 

Mr. GEHRIG. Why was the tank temperature not identified on the test 

specifications or checkout sheets as an item to be observed during 

ground testing at the Kennedy Space Center 1 

Mr. CORTRIGHT. I do not know the answer to that. 
83 

Mr. GEHRIG. Well, the tank temperature was monitored at the Control Center and it was off normal but nobody noticed it. Why do they 

record it if they do not monitor it? 

Dr. PETROSE. May I attempt to answer that? The tank temperature 

TTRS being observed on a scope. It was being recorded on tape. When I 

use the words observed on scope, he is reading a set of numbers and it 

was observed to go up from the range of about minus 300 to plus 80. 

In the judgment of the people observing that rising temperature, 

when it hit 80 and then leveled-did not go any higher-they assumed 

it was at a safe temperature. That was incorrect because that particular gage, or that temperature sensor. has a maximum limit in the 80" 

range. Here was a sensor in effect pegged out but is reading 80. We now 

know the temperature was higher than that. 

The temperature sensor, I might add, is not directly on this heating 

element. It is slightly displaced. It is on the quantity probe but even 

so, the trace will show temperature increase. As TTe plotted later, it 

went from minus 300 to plus 80 in about 3 hours and then leveled off 

at plus 80 but that leveling off was in the temperature sensor. It is the 

upper limits of the gage. The actual temperature was higher. 

Mr. GEHRIG. Somewhere in the report, Mr. Cortright, it seems to me 

I remember reading that there is an alarm, a local alarm on this temperature reading and that nobody knows whether or not the alarm 

vent off or not. Am I correct in that ? 

Mr. CORTRIGHT. What you may be recalling is reference to an alarm 

in Mission Control CenterMr. GEHRIG. That is right. 

Mr. CORTRIGHT. When the pressures of the tank went above limit in 

the short period before the tank ruptured. There was a light that in 

all probability came onMr. GEHRIG. That is what I remember. 

Mr. CORTRIGHT. And indicated that and then went off again when 

the tank ruptured and the pressure dropped down. This was not 

noticed. 

Mr. GDHRIG. I see. 

AEROSPACE SilFETY ADVISORY PANEL 

Dr. Harrington, who are the members of the Aerospace Safety 

Advisory Panel ? 

Dr. HARRINGTON. At the present time we have seven members, with 

myself as Chairman. Mr. Frank DiLuzio, Gen. Carroll Dunn, Dr. 

Henry Reining, Dr. John Hornheck, Dr. Harold Sgnew, and Mr. 

Bruce Lundin. I think I have named all seven. 

Mr. GEHRIG. Would you provide for the record, a brief biography 

of the members of the Panel, please? 

Dr. HARRINGTON. Yes ; I will provide that for the record. 

(The material submitted for the record follows :) 

DR. HAROLD M. AQNEW 

(Weapons Division Leader, University of California, Los Alamos Scientific 

Laboratory, Los Slamos, N. Mex.) 

Birthplace and Date : Denver, Colorado, March 28,1921. 

Educational Background : BA. University of Denrer, 1942 ; MS, University of 

Chicago, 1948 ; PhD, (Physics), University of Chicago, 1949. 
84 

Career Highlights : Employment with the Metallurgical Laboratory, University of Chicago from 1942-1943; joined the Los Alanios Scientific Laboratory as a Staff nieniber froni 1943 to 1946 and 1949 to 1950, appointed Assistant to 

Technical Associate Director from 1951 to 1953 and froni 1954 to 1961 was Alternate Division Leader. On leave froni the Unirersity of California, Los 

Alamos Scientific Laboratory from 1962 to 1964 to become the Scientific Advisor, 

Supreme Allied Commander, Europe. Returned to Los Alamos Scientific Laihratory as Weapons Division Leader. Traveled in Europe, Japan, Korea, Taiwan, 

and Philippines. Member of CP-1 group (first nuclear chain reaction, Chicago. 

1942). Flew with 509th Bombardment Group to Hiroshinia with first nuclear 

weapon. Received the Ernest Orlando Lawrence Award in 1966. 

Memberships : Chairman, Army Scientific Advisory Panel, 1965-pre.wnt ; Member, Aircraft Panel, President's Scientific Advisory Committee, 1966-present ; 

Member, Defense Science Board, 19Wpresent ; Member, NASA Aerospace Safety 

Advisory Panel, 1968-present ; Chairman, U.S. Army Combat Derelopments Command Scientific Advisory Group, 1965 ; Member, U.S. Air Force Scientific Adrisory Board, 1957-196s ; Member, USAF Minuteman Planning Committee, 1961 ; 

Member, Von Karnian Study Group, 1960 ; Consultant to the U.S. Army in 1944 ; 

Fellow, American Physical Society ; R'ew Mexico State Senate, 1955-1961; Los 

dlalli0S Board of Educational Trustees, 1950-1956 (President, 1955) ; Chairman, R'ew Mexico Senate Corporation Committee, 1957-1961 ; Secretary, New 

Mexico Legislative Council, 1957-1961 ; Neniber, Governor's Radiation Advisory 

Council, 1959-1961 ; Phi Beta Kappa : Sigma Xi ; Omicron Delta Kappa. 

Will become Director, Los Alamos Scientific Laboratory on Septemkr 1, 1970. 

PERSONAL HISTORY RESUME-FRANK C. DI LUZIO 

Date of Birth : September 2,1914, Rome, Italy. 

Marital Status : Married. two children. 

Height : 5'6". 

Weight: 150. 

Education: B.S. (CE), Fenn College, ClereLand, Ohio 1935, Case Institute of 

Technology, Cleveland, Ohio ; Cleveland Institute of Technology. 

Additional Training : War Department, Washington, D.C. ; War BIanpower 

Commission. Contract Negotiation and Renegotiation School, Property Management, Property Disposal and Evaluation, Harvard AMP 32 Graduate School 

of Business Administration, Cambridge, Mass., AEC sponsored. 

Experience : .January 1, 197&Present.-President Ecological Systenis Analysis Corporation. The pwposes for which this corporation is formed are to study 

and prepare ecological systems analyses and reconimend courses of acltion ; to 

prevent or abate pollution of air, water and land resources ; to recommend institutional changes and funding methods to implement recommended courses of 

action ; to provide management services to assist in setting up institutions and 

operating procedures to manage anti pollution facilities and services, and to 

engage in any other legal activity. 

January 1, 19?0--Present.-Consultant to E.G. & G. on Environment, Ecology 

and governnient relations. 

January 2, 1968-January 1, 19W-Vice President, EG&G, 1nc.-EG&G is a multi-element company with broad technical capabilities and business interests. 

Research and development, instrumentation systeni design, engineering service 

work and manufacturing, encompass the fields of physics, electronics, optics, the 

nuclear sciences, scientific photography, oceanography and geophysics, the en- vironmental sciences, and computer sciences. In the nonscientific business areas, 

EG&G companies are engaged in the fabrication of standard and custom metal 

products and in largescale scientific nuclear test support service and peaceful 

applications of nuclear energy. 

President, Reynolds Electrical & Engineering Co., Inc-An EGCG subsidiary, 

REECo is the support services contractor to the Atomic Engery Commission at 

the R'evada Test Site and is responsible for base construction, housing and feeding, utilities, tunneling and mining, large-hole drilling, medical services, transportation, architect-engineering, etc., and employs approximately 5,000 people. 

The commercial electrical construction division headquartered in Tempe, Arizona, 

consists of three regional offices across the United States with projects in 37 of 

the 50 states, as well as Guam, Puerto Rico, and Japan. 
85 

~~~~~~~t 1966 to January 2, lg68.-Assistant SeWetaW for Water Pollution 

control, I;.s. ~~~~~tment of the Interior-Acted for the Secretary of the Interior 

in coordinatirlg the rlational pollution efforts under the water Polllltion acts Of 

1965 1966 ; administered Executire Order 11288 which Places in the Secretam 

the responsibility for insuring that federal activities comply with the intent of 

the nater pollution acts : implemented the establishment of state standards in 

,,olllp~iance lyith the water pollution act of 1966 as amended in 1966, and developed criteria for establishing state water quality standards. In addition, retained 

the Secretary’s responsibility for the Office of Saline water, and performed such 

other functions as the Secretary assigned. Participated in and directed Pollutioll 

pontrol studies for Italy, West Germany, France, Israel, Japan, England, and JIexico ; was Department of the Interior coordinator and representative for the 

Iiiteragency committees on air, land, and solid wastes. 

danzlary 1965 to August 1966.-Director, Once of Saline Water, U.S. Department of the Interior-Completed formulation of and placed into effect an aggressive and accelerated program for economically feasible means of desalting sea and brackish waters. Instituted short, medium, and long-range programs leading 

to increased emphasis in engineering for practical application of desalting 

techniques. Participated in and directed water resource development programs 

including surface and ground water development, desalting and waste water 

recovery for Saudi Arabia, Egypt, Iran, Israel, Spain and Italy at the request of 

the U.S. State Department. Implemented the administration’s commitment to 

ulake U.S. water resource technology available to all nations of the world. 

April 20, 1963 to January 25, 1!%5.-Staff Director, Committee on Aeronautical 

und Space Sciences, U.S. Senate-Responsible for preparing program papers for 

the Committee and making special studies for the various committees of the U.S. 

Senate involved in or interested in both the U.S. Air Force and NL4SA space 

programs ; made recommendations to the Senate Committee on Xational Space 

Budgets. 

October 15, 1962 to April 20, 196S.--Vice President & Director, Fairbanks, 

Morse & Co., Hydraulic and Special Projects Division-In charge of engineering 

design, applications egineering, sales, and project administration. Responsible 

for the management and coordination of English Electric and Vickers-Armstrong, 

London, England, license agreements covering their hydraulic products ; served as member of R6;D committee, and management committees of both Fairbanks, 

Morse 6; Co.. and Fairbanks TVhitney Corporation. 

Octobw 1, 1961 to October 15, 1962.--Vice President, Engineering, Fairbanks, 

Morse & Go., Beloit Division.-In charge of all product and special engineering. 

reporting to rice President Group Executive. Products consisted of motors, 

diesel engines, pumps, magnetos, compressors, etc. Engineering organization 

had a staff of approximately 326 profesional and sub-professional people. 

January 1, 1961 to October 1, 1961.-General Manager, Albziqzterqzie Researcli 

C’eizter, Fairbanks, Jl orse & Co.-Reported to Vice President-Government Products. In addition, was also assigned as Assistant for Research and Developrllent 

to President of Fairbanks, Morse & Co. Coordinated research and dereloprnent 

corporation-wide, served as Chairman of R6;D Committee, and eraluated engineering proposals related to national defense programs. 

Llfay 19, 1957 to Janzcary 1, 1961.-Dep?ity Managw, S7bztqlterqrte 0peratio)is 

Oflice, V.S. Atomic Energy CmnzZssion-Assisted Manager in performance of 

executire duties covering all functions assigned to A1;O. Directly responsible for 

the evaluation, coordination. and appropriate action regarding perfornlance of 

functions assigneC1 to the operating divisions in the areas of research, dereloprllent* nlanllfactUring. and quality assurance on atomic ordnance, Directed 

related to storage operations, nuclear materials, management and 

qecuritr administration. Upon special assignment from the -4tomic E~~~~ 

Commiqsion. wa*ington. D.C., executed programs for -%EC Headquarters organizations other than the Division of Military Application. Coordinated for 

the -IEC Division of MilitaQ Application, the exchange of weapon techolm 

the rnited States and the United Kingdom Stomic Energy establish- ments at Aldermaston and Harwell, England. 

Ja?lztary 15. 1.956 to lTfay 19, 1957.-~ssistant Jfanager for ~~anq~fact~crillQ, -4Zb11~e~le Operations Once, U.S. Atomic Energy Commi,ysion-nirected and 

coordinated Aklbwiierque Operations Office development and production complex 

in design, development, and manufacture of nuclear and non-nuclear n7eaponq 

and weapons components. Developed from broad DMA directives detailed ALO 
86 

directives and planning schedules, allocating responsibility and establishing 

delivery requirements, time schedules, and procurement authorizations for the 

several contractors involved. Coordinated plans for all major changes in existing or new development or manufacturing facilities and equipment resulting from 

mission or programmatic changes. Planned and directed the execujtion of ALO 

programs and policies designed to effectively coordinate design activities with 

the manufacturing processes ; planned and coordinated program reporting needs 

of the Manager and the preparation of periodic consolidated progress summaries 

and program statistics. Coordinated weapons development, testing and production with Armed Forces Special Weapons Command-Defense Atomic Support 

Agency for military weapons input. 

August 1952 to January 1956.-Mamager, Los Alanbos Area Oflice, U.S. Atonaic 

Energy Commission-Administered assigned programs in the field of research 

and development of atomic weapons and in this capacity administered dEC’s 

contract with the Los Blanios Scientific Laboratory. Planned and executed a comprehensive program for construction and maintenance of technical and 

other project facilities. Provided supervision of all contract operations, including 

architect-engineering and construction and supply contracts. 

April 1950 to August 1952.-Director, Contmunity Management Division, U.S. 

Atomio Energy Cwmntission, Los Almos-Participated as a member of Area 

Manager’s staff in the continuous review of the assigned Los Alanios Office 

program. 

September 1940 to October 1941.-Engineer, U.S. Bureazc of Reclanlation, 

Parker Dam, California-Field inspection of power plant (one shift), checked 

field changes in construction, steel setting, construction details and procedures. 

March 1958 to September 1940.4r. Engineer, U.S. Bureau of Recluviatio+tAssigned to Coulee Dam, Washington. Field engineer inspection of placing of 

concrete, setting reinforcing steel, and installation of power plant equipment and 

drum gate mechanism. 

PROFESSIONAL ORGANIZATIONAL MEMBERSHIP 

International Association for Hydraulic Research ; Association for Applied 

Solar Energy, Scientific Member, Tempe, Arizona ; National Society of Professional Engineers ; Professional Engineer, New Mexico, Registration So. 3262 ; 

Professional Engineer, Alaska, Registration No. 859-E ; Professional Engineer, 

Nevada 2969 ; American Geophysical Union, Sational Academy of Sciences, 

Washington, D. C. ; Seisniological Society of America ; American Society of Civil 

Engineers. 

President’s Commission on M arime Hcience, Engineering and Resoitrces 

Presidential appointment, January 1967. Established by the June 1966 Marine 

Resources and Engineering Development Act to examine the Sation’s stake in 

the development, utilization, and preservation of our marine environment ; to 

review all current and contemplated marine activities and to assess their adequacy to achieve the national goals set forth in the Act ; to forniulate a comprehensive, long-term, national program for marine affairs designed to meet present 

and future national needs in the most effective possible manner ; to recommend a plan of government organization best adapted to the support of the program. 

NASA Aerospace Safety Advisory Panel 

Congressional Appointment, May 1968. Established by the National Aeronautics and Space Administration Authorization Act, 1968, the Panel reriews 

safety studies and operations plans referred to it and makes reports thereon ; 

advises the Administrator with respect to the hazards of proposed or existing 

facilities and proposed space missions and program operations and with respect 

to the adequacy of proposed or existing mission modes and safety standards; 

and performs such other duties as the Administrator may require. 

National Water Commission 

Presidential Appointment, October 1968. Established by the Kational Water 

Commission Act, Congress has directnd the Commission to review present and 

SELECTIVE MEMBERSHIP 
87 

anticipated national water resource problems, assess the nation’s future water 

needs and identify alternative ways to meet them; to give consideration to 

conservation, more efficient use of existing supplies, reduction of pollution, innovations to encourage the highest econonlic use of water, interbasin transfers, 

and technological advances. The technology includes but is not limited to desalting, weather modification and purification and reuse of waste water. Congress has directed that the Commission consider not only the best technology, 

but also the economic, social, and aesthetic consequences of water resource development. 

;\IAJ. GEN. CARROLL HILTON DUNN 

Carroll H. Dunn was born in Lake Village, Arkansas August 11, 1916. He 

graduated from the University of Illinois, Urbana, Illinois, with a Bachelor of 

Science degree (i\lechanical Engineering) in 1938. His military career began 

July 1938 \>Then he Tvas ,commissioned a Second Lieutenant in the Regular Army 

by professional examination. His first assignment was at Laredo, Texas with the 

Eighth Engineer Squadron, First Cavalry Division. In February 1941 he was assigned to the Engineer Replacement Training Center, Fort Leonard Wood, 

Missouri. 

In November 1942 he was assigned to the 30th Infantry Division as Division 

Engineer and concurrently Commanding Officer, 105th Engineer Combat Battalion. He remained assigned to this Division through training at Camp Blanding, Florida ; Oamp Forest, Tennessee ; Camp Atterbury, Indiana, and moved overseas with the Division to England. From June 1944 to May 1946 he participated n-ith the 30th Division in combat in Europe, going from the Omaha 

Beach to Magdaburg, Germany. During this period he was wounded by an enemy 

mine during the attack on Saint Lo, and spent two months in the hospital in 

England, returning to join his unit for the final drive into Germany. 

In July 1945 he mas assigned to the Second Infantry Division as Assistant 

Chief of Staff, G-4, a position which he held until May 1946, when he mas ordered 

by the Army to the State University of Iowa, Iowa City, Iowa, to take graduate 

engineering study. He graduated in June 1947 with a Master of Science degree 

in Ciril Engineering, following which he was assigned to the Engineer School, 

Fort Belvoir, Virginia, as Instructor in combat engineer activities. 

From October 1949 to August 1952 he was assigned to the Engineer Section 

GHQ Far East Command, where his principal duties concerned staff supervision 

of the construction activities of that Command. 

He returned to the United States in August 1952 and was assigned as Director, 

U.S. Army Waterways Experiment Station, Vicksburg, Mississippi, a principal 

research activity of the Corps of Engineers in the field of hydraulics, soils, and 

concrete. He continued in this assignment until July 1955, when he was ordered 

to Washington, D.C. to become Esecutive Officer to the Chief of Engineers, United States Army. an assignment held until August 1958, when he was selected 

for attmdance at the Industrial College of the Armed Forces, Washington, D.C. 

Upon graduation from the Industrial College. he was assigned to Thule, Greenland as Area Engineer and was responsible for construction of facilities for the 

Sation’s first Ballistic Missile Early Warning Sptem. Upon return to the United 

States in July 1960, he was assigned to the newly organized Corps of Engineers 

Ballistic JIissile Construction Offiice at Los Angeles, Jvith the dual position of 

Deputy Commander and Director of Titan I1 Missile System construction. 

With his nomination for ~)roniotion to Erigadier General on 18 January 1962. 

he was reassigned as Division Engineer, r.S. Army Engineer Division, Southwestern. Dallas, Texas. effective 1 March 1962. In this assignment he lvas re- sponsible for a construction progrml exceeding 300 lllillion dollars per year. 

-imong the many construction projects under his supervision lvere the IIanned 

Spacecraft Center at Houston and the 1.2 billion dollar prograln to improve the 

Arkansas River for navigation, flood control. water supp1y. and power. 

On 1 -kugUst 1964 he was assigned to the Eighth U.S. Army in Korea as Deputy Chief of Staff. He held that assignment until 17 January 1966, when he Jvas re- 
88 

assigned to Vietnam to assume directive control of all Department of Defense 

construction programs in Vietnam. He served as Director of Construction, United 

States Military Assistance Command, Vietnam, until 30 June 1966. 

On 1 July he was reassigned as Assistant Chief of Staff for Logistics (J4), 

United States Military Assistance Command, Vietnam. In this assignment, he was responsible for coordination of all logistics support for the U.S. and Free 

World Forces in Vietnam, essentially assuring that the material, equipment and 

transportation, needed to support combat operations, were available. He con- tinued to hold this position until 15 September 1967. 

On 16 October 1967 he was assigned as Director of Military Construction, 

Office of the Chief of Engineers, Washington, D.C. In this position he was re- sponsible for military construction within the Army, and for construction and 

design work performed for the Air Force and ;he Sational Aeronautics and 

Space Administration, and other Government Agencies as assigned. He was also responsible for the Army Nuclear Power Program and specialized fallout 

shelter engineering support for Civil Defense. The dollar value of work assigned 

to the directorate totals approximately one billion a year. 

General Dunn was appointed Deputy Chief of Engineers on 1 August 1969. 

PERSONAL DATA : 

Born : August 11,1916, Lake Village, Arkansas. 

Father: William L. Dunn, Sr., Mother: Ruth Dewey Dunn. Both reside in 

Lake Village, Arkansas. 

Married Letha E. Jantz 11 November 1939 at Moline, Illinois. 

Children: Carolyn J. (hlrs. Douglas L. Caldwell, Apt T-1, Seminary Forert 

Apts., Moo No. Pickett Street, Alexandria, Virginia 22304). 

Capt. Carroll Hilton, Jr., Quarters 5528, Pope Road, Ft. Belroir, Va. 23060. 

Official Address : %The Adjutant General, Department of the Army, Washington, D.C. 20310. 

EDUCATIOR' 

University of Illinois (BS in Mechanical Engineering)-l938 ; Command and 

General Staff School, Seventh General Staff Class-1942 ; State University of 

Iowa (MS in Civil Engineering)-l947; Industrial College of the Armed 

Forces-1959. 

CHRONOLOGICAL LIST OF PROMOTIONS 

Rank Temporary Permanent 

(AW (RA) 

1 DOR July 1,1961. 1 DOR Jan. 29,1966. 3 DOR Nov. 25,1967. 
CHRONOLOGICAL LIST OF ASSIGNMENTS 

Assignments From To 

8th Engineer Squadron, Fort Mclntoch, Tex ______.______________..__.

July 1938.. .____ Feb. 1941. 

Engineer Replacement Training Center, Ft. Leonard Wood, Mo ._______________._._ Feb. 1941 ______ ~ July 1942. 

303d Engineer Combat Battalion Camp Butler N.C ._____________._______________ July 1942. ______ Nov. 1942. 

Division engineer, 30th Infantry’ Division and’commanding officer, 105th Engineer Combat Battalion, Camp Blandlng, Fla., Camp Forest, Tenn., Camp Atterbury, Ind England and Continental Europe ___...______________________

Nov. 1942 _____. May 1945. 

Commhding ofher 1153d Engineer Combat Grou Le Havre France ________ _____ May 1945 _..___ ~ July 1945. 

Assistant chief of stiff, G4,Zd Infantry Division, kmpswift: Tex. and Fort Lewis, 

Wash ... .................... __ ____________________--.-.-.-....----.

July 1945 _._____ May 1946. 

Graduate student, State University of Iowa, Iowa City, Iowa _______________.._____ May 1946 __.._._ June 1947. 

Engineer school instructor, Fort Belvoir, Va ____________._.__.__-..______July 1947.. _..__ July 1949. 

Engineer section GH Far East Command, Tokyo, Japan ___________________._Sep. 1949 _______ Aug. 1952. 

Director, U.S. Army iaterways Experiment Station, Vicksburg, Miss _._________..__ Sep. 1952 _..._._ June 1955. 

Executive officer to the chief of engineers, U.S. Army, Washington, 0.C ____________ July 1955 ___..._ Aug. 1958. 

Student, Industrial College of the Armed Forces, Washington, D.C ____.___....__.._ Aug. 1958 ______ June 1959. 

Area engineer, Thule, Greenland _____________________________July 1959 .____._ July 1960. 

Deputy commander, Corps of Engineers Ballistic Missile Construction Office, and 

Director, Titan II construction _____..________.____-----.-.-.-.-.---.-..-.

~ Aug. 1960 ______ Feb. 1962. 

Division engineer, US. Army Engineer Division, Southwestern, Dallas, Tex. ________ Mar. 1962 ______ July 1964. 

Deputy Chief of Staff, 8 U.S. Armh,Seoul, Korea ~ __._ ~ _________ Aug. 1964 ______ Jan. 1966. 

Director of construction, US. Wary Assistance Command, Vietnam, Saigon, Vietnam _______________________________________

Feb. 1966 _____._ June 1966. 

Assistant Chief of Staff for Logistics, J-4, U.S. Military Assistance Command, Vietnam, Saigon, Vietnam ______.____________________________

July 1966 _____._ Se . 1967. 

Director of military construction,Office of the Chief of Engineers, Washington, D.C.- Oct. 1967 ....... Jub 1969. 

Deputy Chief of Engineers, Washington, O.C.-.. _________ ~ _____________ ~ ________ Aug. 1969 ______ 

LIST OF CITATIONS AND DECOBATIONE 

Distinguished Service Medal; Silver Star; Legion of Merit; Bronze Star 

Medal with two oak leaf clusters and “V” device ; Army Commendation Medal 

with two oak leaf clusters; Air Force Commendation Medal with one oak leaf 

cluster ; Purple Heart ; French Croix de Guerre are Palm ; Belgian Fourragere. 

EEBVIOE MEDU 

American Defense Service Medal ; American Campaign Medal ; World War I1 

Victory Medal ; nurope-Middle East Campaign Medal with five campaign stars ; 

Ko‘rean Service Medal; UN Service Medal; Army of Occupation Medal, Germany ; Army of Occupation Medal, Japan ; National Defense Service Medal ; 

Vietnam Service Medal with three campaign stars. 

PEBSORAL BAOKQBOUND MATEBIAL 

Interests and Hobbies.-Golf ; Bowling ; Spectator Sports ; Photography (Color 

Slide). 

CIVIO AND PROFESSIONAL ACl’IVITIES 

Registered Professional Engineer (District of Columbia and Texas) ; Fellow, 

American Society of Civil Engineers ; Member, Society of American Military Engineers ; Association of the U.S. Army ; Member, NASA Aerospace Safety Advisory Panel. 

BELIQIOR 

Protestant (Baptist) : Deacon, Sunday School Teacher ; Active in Church in area to which assigned. 

CHARTJN D. HABBINQTON 

Dr. Harrington was born July 22, 1910. He received his B.S. degree in Chemistry from Harvard College in 1937 and M.A. and Ph.D. degrees in Chemistry 

from Harvard University Graduate School in 1939 and 1941, respectively. 

He joined the Mallinckrodt Chemical Works in 1941 as a Research Chemist 

and was assigned to work with the Manhattan Project in 1942 where he assisted 

in the development of the initial process used for uranium purification. In 1944 

he became the Technical Director of Mallinckrodt’s Uranium Division and in 
90 

1952 was appointed Manager of that division with responsibility for the direction of activities of the Destrehan Plant. In 1968 responsibility for the then new Weldon Spring Plant was assumed. Both of these plants were operated by 

Mallinckrodt for the Atomic Energy Commission. In 1960 he was elected a Vice 

President of Mallinckrodt and in May 1961, when the Nuclear Division of 

Mallinckrodt merged with the Nuclear Fuels Division of Olin and the Nuclear 

Development Corporation of America to form the United Nuclear Corporation, he 

became Vice President, Chemical Division and a member of the Board of Directors of the new corporation. In 1963 he became the Senior Vice President of 

United Nuclear Corporation. 

In July 1965, Douglas United Nuclear, Inc. was formed as a subsidiary owned 

jointly by the L)ou:las Aircraft Comjiany (now McUonnell Douglas Corporation) and United Nuclear Corporation, and Dr. Harrington became the President and General Manager of this new corporation. Douglas United Nuclear, 

Inc. was formed to carry out the joint activities of the two parent companies in 

the Tri-Cities, Washington area. These activities include the management of the 

production reactors and fuels fabrication facilities at Hanford for the Atomic 

Energy Commission and the furtherance of other commercial objectives in the 

nuclear energy field. As a complement to the development of the Tri-Cities area and as a base of commercial activities on the part of the several new contractor 

organizations engaged at Hanford, Douglas United Nuclear has been actively 

participating in the expansion of the Graduate Center in the area. In addition to his current position as President and General Manager of 

Douglas United Nuclear, Dr. Harrington serves as a Director for the corporation and also as a Director and Vice President of United Nuclear Corporation. 

He is also a Director of the Tri-City Nuclear Industrial Council. He has served as a member of the Atomic Energy Labor-Management Advisory Council since 

its inception and in February 1968, was appointed to a six-year term as a member of the Aerospace Safety Advisory Panel, National Aeronautics and 

Space Administration, and is currently serving as Chairman of this Panel. 

Dr. Harrington is co-author of “Uranium Production Technology” which was published in 1959 and has also published or delivered a number of papers, including those delivered at the Paris (1957) and Rome (1963) Conferences on Nuclear Energy. He was recipient of the 1960 Mid-West Award of the 

American Chemical Society for contributions to technology in the nuclear 

energy field. 

Dr. Harrington is a member of the Bonneville Regional Advisory Council 

and a member of the Advisory Board, College of Engineering, Washington 

State University. He is a fellow of the American Nuclear Society and a member 

of the American Chemical Society, the National Space Club, the American 

Association for the Advancement of Science, the American Management Association and the Atomic Industrial Forum. 

JOHN AUSTIN HORNBECK 

J. A. Hornbeck is a vice president of the Western Electric Company and 

president of the Sandia Corporation, a nonprofit Western Electric subsidiary. 

Sandia Corporation operates Sandia Laboratory at Albuquerque, N.M., Liver- more Laboratory at Livermore, Calif., and other smaller facilities for the Atomic 

Energy Commission. He is also a director and member of the executive com- mittee on Sandia Corporation. His office is in Albuquerque. 

He was born on November 4, 1918 in Northfield, Minn,, and graduated from 

Central High School in Kalamazoo, Mich., in 1935. He received a Bachelor of 

Arts degree in physics from Oberlin College in Ohio in 1939 and a Doctorate 

in Physics from the Massachusetts Institute of Technology in 1946. While in 

college, he was a member of the soccer and debating teams and was captain 

of the tennis team during his senior year. 

Following his graduation from Oberlin, he was a teaching fellow at X1.T. 

for two years before joining the National Defense Research Committee as a technical aide. He returned to M.I.T. as a research assistant in 1943, becoming 

secretary of the N.D.R.C.’s Land and Mines Committee a year later. 

Mr. Hornbeck began his Bell System career in 1946 as a research physicist 

in the physical electronics department of Bell Telephone Laboratories. Transferred to the transistor research department in 1951, he was placed in charge 

of the semiconductor physics department in 1952, becoming head of the solid state 

devices department the following year. 
91 

From 1956 to 1958, he was director of electron device development and from 

1958 to 1962. he was executive director of the Semiconductor Device & Electron 

Tube Division. In 1962, he was named president and a director of Bellcomm, 

Inc. (jointly owned by L4merican Telephone and Telegraph Company and 

Western Electric and located in Washington, D.C.). 

JIr. Hornbeck was elected vice president of Western Electric Company, effective September 1, 1966. At that time, he also became vice president, a director, 

and a member of the executive committee of Sandia Corporation. He assumed 

his present position on October 1, 1966. 

A fellow in the Institute of Electrical and Electronic Engineers and the -4merican Physical Society. he is a member of the -4merican Bssociation for the Aavancement of Science, Phi Beta Kappa, Sigma Xi, Delta Sigma Rho, the Cosmm 

Club (Washington, D.C.), and the Four Hills Country Club (Albuquerque, S.M. ). 

He is a member of the Aerospace Safety Advisory Panel to SBSA and serves as chairman of the Advisory Board, Sational Bureau of Standards Institute of 

Basic Standards. He has contributed articles to Phusical Review and other professional journals. 

RIr. Hornbeck is president of the Albuquerque United Community Fund and a director of the Albuquerque Symphony .. Orchestra and the Presbyterian Hospital 

Center Foundation. 

He married Eniily Elizabeth Aldrich on January 31, 1942 in Wauwatosa, Wise. 

They have two daughters- (hIrs.) Joan Aldrich Smith and Deborah Snn-and 

three sons-Kirk dustin, John Frederick, and Christopher Wolfe-and live at 

1516 Sagebrush Trail, S.E., Albuquerque 87123. 

his leisure time. 

Rfr. Hornbeck is active in Boy Scout activities and enjoys golf and bridge in 

BRUCE T. LUNDIX, DIRECTOR, LEWIS RESEARCH CEKTER, 

SATIONAL AERONAUTICS AND SPACE ,4DMINISTRATIOS 

Bruce T. Lundin, a native of Alameda, California and a graduate in Rlechanical Engineering of the University of California in 1942, began his scientific ca- reer with the Kational Advisory Committee for Aeronautics in 1913 at the Lewis 

Research Center. He was initially engaged in heat transfer investigations and 

in improving the performance of our World War I1 aircraft engines. In 1946 he was placed in charge of the Jet Propulsion Research Section, which conducted some of this country’s early research on turbojet engines. 

In 1952 RIr. Lundin was appointed Chief of the Engine Research Division at 

the Center, and became responsible for the full-scale engine program. Work that 

he directed over the nest several years contributed significantly to the performance and reliability of today’s commercial transport and supersonic aircraft 

jet engines. He also pioneered in research on large-scale ramjet engines. 

When R’,4CA4 became the nucleus of the present National Seronautics and 

Space ddministration in October of 1958, Rfr. Lundin was appointed an iissistant 

Director of the Center. In this capacity he directed much of the Center’s es- panded role in space propulsion and power generation. This role was further en- larged in December, 1961 when he ms appointed Associate Director for Development. The responsibilities emb~aced development of turbojet engines, chemical 

rockets, electric thrusters for spacecraft propulsion, and electric pan-er generating systems for spacecraft using chemical, qolar and nuclear energy sources. He 

also directed the development and operation of SSSh’s Centaur and Agena 

launch vehicles for unmanned spacecraft and of spacecraft for investigating adranced methods of space propulsion. 

In >fay 1968 Mr. Lundin went to NASA Headquarters as Deputy Associate 

Administrator for the Office of Bdvanced Research and Technology, and in Bfarch 

1969 was named Acting Associate Administrator. On November 1, 1969 he was appointed Director of the Lewis Research Center. 

Nr. Lundin is a member of Tau Beta Pi and Sigma Xi, an Associate Fellonof the American Institute of Aeronautics and Astronautics, and a Fellow of the 

Royal Aeronautical Society. He is also a member of several governmental advisory committees, including the NASA Aerospace Safety Advisory Panel, the 

Scientific Advisory Board of the U.S. Air Force, and the NASA Research Advisory Committee on Space Vehicles. In 1965, he received the NASA Medal for 

Outstanding Leadership. 
92 

HENRY BININU, JE. 

Henry Reining, Jr. became the first dean of the newly constituted Von KleinSmid Center for International and Public Affairs in 1967. His administration 

includes three instructional components-the School of Politics and International 

Relations, with two departments, i.e. Political Science and International Relations, the School of Public Administration, and the Graduate Program of Urban 

and Regional Planning. His role as dean provides him with the opportunity to 

bring to bear all of the resources of the University which are devoted to gov- ernmental affairs. 

A former president of the American Society for Public Administration, he has 

served government at all levels. He pioneered in the development of the A’ational 

Institute of Public Affairs in Washington, D.C., serving as its first executive 

director. His activities at the national level have also included membership on a number of committees and commissions. He is presently a member of the 

National Aeronautics and Space Administration’s Aerospace Safety Advisory Panel. 

During World War 11, he served as consultant to the United States National 

Resources Planning Board in the establishment of the National Roster of Scientific and Specialized Personnel, to the War Department, School of Administration, 

Port Washington, and to the United States Civil Service Commission for whom 

he conducted the first generalist examination for executives and administrators 

(in 1940) and assisted in the establishment of the Public Administration Examining Division. In 1943-44, he traveled to Brazil for the United States Coordinator of Inter-American Affairs to assist the Brazilian National Department of Administration. 

He has been a consultant to the United States Commissioner of Indian Affairs on administrative matters, and mas instrumental in the selection and training 

of superintendents of Indian reservations. He was consultant to the Secretary of the Interior for whom he made a survey of Boulder City, Nevada, which 

ultimately enabled it to become a self-governing community. He was a member 

of the U.S. Civil Service Commission’s Regional Loyalty Board for a number of 

years, and of the national Loyalty Review Board. 

He went to Brazil again in 1951-52, for the United Nations, to chair a multinational faculty team to set up South America’s first School of Public Administration in Rio de Janeiro. In 1953, he journeyed to Turkey to represent the U.S.A. 

at an international congress at Istanbul, and to advise on the development of a regional public administration institute at the University of Ankara. 

In 1954, Dean Reining made the survey in Iran which led during the next seven years to a joint effort with the University of Tehran to establish an Institute of Public and Business Administration. now a permanent institution. 

During the summer of 19.54, he went to the Philippines to assist the University of the Philippines in the development of its Institute of Public Administration, 

especially its in-service training program. In 1957, he surveyed the training needs of the government of Pakistan, which 

effort led to the setting up of an executive development program at USC for 

groups of top civil servants, totaling almost 100 during the succeeding three year period. 

Dean Reining moved the USC Pakistan program to Pakistan in 1960. At that 

time, he assisted in setting up a system of three institutes of public administration in two wings and at the center of that Government. He also assisted in 

establishing the curricula at two universities, the IJniversity of Dacca and the 

University of Punjab. 

Dean Heinine has also been active in state and lwal matters. He was a con- sultant to the California State Assembly in 1950, and made a study of the reorganization of the State Government. He was, during 1957-58, Chairman of the 

Los Angeles County Charter Commission, which presented a thorough revision 

to the County’s Board of Supervisors. He is presently Chairman of the Los 

Angeles City Charter Commission which presented a new charter in July 1969 

and has participated in the City Council’s hearings held since that date. 

Dean Reining began his academic career and his association with USC in 1932, as Assistant Professor, School of Public Administration. 

During the period 1934-1936, he was a member of the Politics faculty at 

Princeton University and a Research Associate, Local Government Surveys ; 

1935-1945, Educational Director, National Institute of Public Affairs in Wash- 
ington. D.C. ; 194346. Management Consultant with the firm of Rogers. Slacle 

ancl Hill, Seiy Tork City; 194647, Assistant Executive Director, Port of Sew 

Torli Authority, Sew Torli City. 

In 1947, Dr. Reining returned to USC as Professor of Political Science alld 

Public Administration. In 1953, he became Dean of the School of Public ddministration and in 196i, Dean of the Von KleinSmicl Center for International and 

Public Affairs. 

His professional associations hare inclucled : American Society for Public Administration (national president. 19.57-55) : Council on Graduate Education for 

Public Administration (chairman, 196466) ; Pi Sigma Alpha (national vicepresident, 106%1960) ; American Political Science Association ; Public Personnel 

Association : Western Governnlental Research Association. (Board of Goreriors. 19~+1960) : et al. 

Dean Reining received his A.B. in Political Science in 1929 at the Univerqity 

of Akron: his A.11. in Politics at Princeton University in 1930, and his Ph.D. 

in Poiltics in 1932. also at Princeton UniversitF. 

Dr. Reining is the author of a number of books. monographs. and articles for 

learned journals. 

He was born September 1.5. 1907. He is married to Darline Diekmann Reining. 

and has four chilclren : William Henry, Judith Ellen, Susan Elisabeth, ancl Richard Charles. 

-\DVISORT I'ASCL -\CTIVITIES 

JIr. GEHRIG. Where did the Panel meet to consider and review the 

work of the Apollo 13 Reriev- Board ? 

Dr. HARRISGTOS. We met at tlie Jlaiiiied Spacecraft Center during 

the time that the Review Board was meeting there. I first Tent there 

at JIr. Cortriglit's invitation on April 29 and 30 to cliscuss the planned 

scope of the jiiqiiiry and to arrange for the Panel's review of the 

procedure-. The Panel then met on May 6 in Houston with both the 

dpollo 13 Review Board aiicl the JISC _kpollo 13 Inr-estigating Teain 

to review the inquiry process, the evidence, and the determinations 

and findings to date. Then v-e met again on May 27 in iyorking sessions 

with the Panel chairmen of the Apollo 13 Eeviex- Board. One of the 

Panel members and the Executire Secretary of the Panel made aclditional trips to rei-ieJT the continuing operations of the Board. 

Mr. GEFIRK~. Hon- iiiaiiy hies did the panel meet to discuss the Revien- Board's ~ork ? 

Ih. I-T*\RRISGTOS. ThePanel met twice n-ith. as I haye just mentioned. at Hoiiitoii. with the Board. Based on these iiieetings together 

v-itli a re\-iew of the Board report it?elf. tlie Panel met on ,June 19 

to prc>!iare the report to tlie Acliiiini~trator. 

JIr. GEIIEIG. Were the findings of tlie Panel iin:iiiiiiioiis or T~S there 

di-ent ? 

Dr. Hmr,Iscn)s. There ~ras no clisent. 

T-nrioiiq imiiliers liad various inputs into tlie report v-hich vere 

:tccelxccl by all the others. 

311.. GEIIRIG. So the report is iiiiaiiiiiious ? 

Dr. H-incIscTros. Ye<. 

311.. GEIIRIG. That vns the Panel charter with respect to the Apollo 

12 accident? 

Dr. H.<.RRTSGTOS. We had a specific directive from the -Iclministrator. Dr. Paine, to review the procedures and findings of the Apollo 

13 Be1 iex- Board aiicl report on those and report on the findings 

to the -1dmini~trator. 

JIr. GEIIEIG. 7Toulcl you have that placed in the record, please ? 

1)r. Ii.mxsc7~ros. Yes. 

(Tlic material subiiiittecl for the record follom :) 

47-47C-7ll--i 
94 

Na4TIOS-4L AEROXA4UTICS ASD SPlCE A~DMIXISTRATIOS. 

OFFICE OF THE ADJIINIGTRATOR. 

Walshington, D.C., April 20, 1910. 

To : Dr. Charles D. Harrington, Chairman, Aerospace Safety Advisory Panel. 

Subject : Review of Procedures and Findings of Apollo 13 Review Board. 

Attachment : (a) Jfemoranduni sated Bpril 17, 19i0, to Mr. Edgar M. Cortright, 

subject : Establishment of Apollo 13 Review Board. 

References : (a) Section 6 National Aeronautics and Space Administration Authorization Act, 1968 ; (b) NMI 1156.1PAerospace Safety Advisory Panel. 

1. In accordance with References (a) and (b), the Aerospace Safety Advi<or:\- 

Panel (hereafter referred to as the Panel) is requested to revien- the procedures 

and flndings of the Apollo 13 Review Board (hereafter referred to as the Board) 

established by Bttachment (a). 

2. The procedures established by the Board will be made available to the 

Panel for review and coninient as provided in paragraph 4(a) of Attachment (a). 

3. As Chairman of the Panel, you are designated an Observer on the Board. 

In this capacity, yon, or another member of the Panel designated by you. are authorized to be present at those regular meetings of the Board you desired to 

attend. You are also authorized to receive oral progress reports from the Chair- man of the Board or his designee from time to time to enable you to keep the 

Panel fully informed on the work of the Board. 

4. The final report and any interim reports of the Board will be made available 

promptly to the Panel for its review. 

5. The Panel is requested to report to us on the procedures and findings of 

the Board at such times and in such form as you consider appropriate, but no later than 10 days after the submission to us of the final report of the Board. 

T. 0. PAINE, 

GEORGE 31. Low, 

Administrator. 

Deputy Administrator. 

Enclosure. 

S-4TIOSAL -4EROSAUTICs AND SPACE Abl\fINISTRATION, 

OFFICE OF THE ADhlIXISTRA4TOR, 

Washington, D.C., April 17, 1970. 

To : Mr. Edgar M. Cortright. 

Subject : Establishment of Apollo 13 Review Board. 

References : (a) KXI 8621.1-&fission Failure Investigation Policy and Proce1. It is SASA policy as stated in Reference (a) “to inrestigate and docnment the causes of all major mission failures which occur in the conduct of its 

space and aeronautical activities and to take appropriate corrective actions as a result of the findings and recommendations.” 

2. Because of the serious nature of the accident to the Apollo 13 spacecraft 

which jeopardized human life and caused failure of the Apollo 13 lunar mission. 

we herebr establish the Apollo 13 Review Board (hereinafter referred to as the 

Board) and appoint you Chairman. The members of the Board will be qualified 

senior individuals from NBSA and other Government agencies. After consultation with yon, we will : (a) Appoint the members of the Board and make any subsequent changes 

necessary for the effective operation of the Board ; and 

(b) Arrange for timely release of information on the operations, findings, 

and recommendations of the Board to the Congress, and, through the SASA 

Office of Public dffairc. to the public. The Board will report its findings and 

recommendations directly to us. 3. The Board mill : 

(a ) Review the circumstances surrounding the accident to the spacecraft 

JT-hich occurred during the flight of L4pollo 13 and the subsequent flight and 

ground actions talien to recorer, in order to establish the probable cause or causes of the accident and assess the effectiveness of the recovery actions. 

(b) Review all factors relating to the accident and recovery actions the 

Board determines to be significant and relevant, including studies, findings, 

recommendations, and other actions that have been or may be undertaken by 

the program offices, field centers, and contractors involved. 

dures ; (b) KJII 1166.14-Aerospace Safety Advisory Panel. 
Y3 

(c) Direct such further specific investigations as may be necessary. 

(d) Report as soon as possible its findings relating to the cause or causes of the accident and the effectiveness of the flight and ground recovery actions. (e) Develop recommendations for corrective or other actions, based upon its 

findings and determinations or conclusions derived therefrom. 

(f) Document its findings, determinations, and recommendations and submit a final report. 

4. As Chairman of the Board you are delegated the following powers: (a) To establish such procedures for the organization and operation of the 

Board as you find most effective; such procedures shall be part of the Board's 

records. The procedures shall be furnished the Serospace Safety Advisory Panel 

for its review and comment. 

(b) To establish procedures to assure the execution of your responsibilities in your absence. 

(c) To designate such representatives, consultants, experts, liaison officers, 

observers, or other individuals as required to support the activities of the 

Board. You shall define their duties and responsibilities as part of the Board's 

records. 

(d) To keep us advised periodically concerning the orgainzation, procedures, 

operations of the Board and its associated activities. 

5. By separate action we are requesting the Aerospace Safety Advisory Panel 

established by Reference (b) to review both the procedures and findings of 

the Board and submit its independent report to us. 6. By separate action we are directing the Associate Bdministrator for BIanned 

Space Flight to : (a) Assure that all elements of the O5ce of Manned Space Flight cooperate 

fully with the Board and provide records, data, and technical support as requested. (b) Undertake through the regular OMSF organization such review, studies, 

and supporting actions as are required to develop recommendations to 11s on corrective measures to be taken prior to the Apollo 14 mission with respect to 

hardrare, operational procedures, and other aspects of the Apollo program. 

5. All elements of SASA will cooperate with the Board and proTide full 'upport within their areas of responsibility. 

T. 0. PAISE, 

GEORGE 31. Low. 

A dm in i s tra t o r. 

Deputy Administrntor. 

JIr. GEHRIG. Did the Panel hare comnlete freedom to in1-estipate 

anything they naiited with regard to the accident or with regard 

to tl3e 'Tmcecliireq uqed 1~ the Review Board? 

nr. H ~RRTXGTOX. Yes, we did. So jirobleni. 

JIr. GEIIRIG. Prior to the Apollo 18 accident. what did the actirities 

of the Panel coniist of? 

Dr. TT\ncrxc~o~-. The Panel's first year was spent in a surrey of 

the _\i-ollo prorrrain nianamnient swteni and the svstein for hazard 

irlentifimticn. The reT-iew in1 olred ctaff and program elements at 

S_\ S-1 IIeadr~narter~. tlie JInnned Space Flight Centers, and the 

majorit7 of principal contractors for tlie spacecraft. launch vehicle. 

and -imllo iiii+.ioii cunport. The Panel met in qewioii 22 days. 

Nr. Grrrnrc. n'ould yon cay that the Panel v-as well prepared for 

tl7e;r revie~~- of the -4 pollo 13 Board's procedures, then, and to review 

the arcicleiit ? 

Dr. Hinr-rs~~os. Yes. In addition to this general program which 

I have ontlined, we had further assignments from the Administrator 

~uch as a wrie~~ of the inrestigation process involved in the LLRV 

and LLTT accidents reI-iev boards, and a reyiew of the management 

process for the evaluation of risks inherent in reducing Saturn static 

testing and launch operations. 

PRIOR I( II~ TTI~~ or r WLI. 
TECHSICAL XASTAGEXENT 

RIr. GEIIRIG. Dr. Warrington, I hare one more question in two parts 

that I dl put in the record for you to answer, if that is satisfactory. 

On the second page of your letter in the fifth paragraph you say, 

and I pick up the quote : 

. the lack of understanding of thermal switch performance and the informal em in the valuation of the test procedure change reflect more a failure of 

human judgment than a failure in the requirements of the technical management iystem.” 

Khat do you iiiean by that ? 

Further. -1s I unclerstancl it : 

(1) There was no procedure for checking the switch design against 

tlie ground support systein at the Kennedy Space Center (See pp. 5-9 

ancl 5-16 Fincling So. 16). 

(2) The technical inanagenient system did not require that prior 

anomalies 011 a criticd subsystem be brought to the attention of the 

decisioiiinaliers when a significant anomaly occurred in that critical 

subsystem. 

(3) There was 110 procedure requiring the ground support personnel at Kennecly Space Center to inonitor the heater current readings 

on the oxygen tank heater control panel. 

(4) The diniensioning on the oxygen storage tank fuel line was 

sucli that looseness to the point of incomplete connection vas possible 

in the event of worse case tolerance buildup. 

(5) The oxygen tanks contained a combination of oxidized comhustible ninterial and potential ignition sources which was precisely 

the probleiii of tlie A\pollo 204 acciclent. 

All of these seem to iiie to be failures of technical nianageiiient : are 

tliev not I 

(The answer supplied for tlie record follovs :) 

The Cortright Report called for redesign of the sub-system components and 

strengthening of the specific procedures appropriate to the situation and the 

Panel concurred in this. We stated that recommendations made for the re- design of the system must improve the nianufacturability and minimize the 

failure modes of the system under out of tolerance conditions. We also said that 

recommendations for improvements in the analysis and yalidation of non-standard situations are certainly necessary and should be implemented immediately. But to hare said this is not to have said enough. Procedures are the con- sequence of human judgment and human judgment must knowledgeably implement them. 

A series of individual judgments were made during the life cycle of the tank 

without complete awareness of the consequences or their possible interdependence and cumulative effect. If the cumulative situation had been recognized 

the procedures were in existence for an appropriate management response. Lack 

of continuity of engineering evaluation of the subsystem over an extended 

Iieriod of time further compounded the situation. 

Thus, as we noted, the problem is also to sustain human motivation and assure working familiarity with the sub-systems and their hazards. The Panel, as did the Apollo 13 Review Board, recognized the significance of this variable ; 

and we hare further cited some of its implications for the future. 

Mr. GEHRIG. Mr. Myers or Dr. Petrone, I have a number of questions 

that I will pnt in the record for you to answer. 

(Additional questions snbniitted by Mr. Gehrig and answers SUPplied for the record are as follon-s :) 

Qltestio?z 1. In appendix D. p. 92. the Board stated that the comprelbensive 

recicic initiated bu the XSC Apollo 13 incertiuating team of all CSX and LM 

tanks. CalCes, and associated system elements dn wliich oxygen or OXidi#ers are storcd. roil trolled or distribiited should be p~OscCtLted %igOrO?lslll. What is the statiis of thin wviczo? 













Page 

13 

i2 

4.5 
58 

81 

9 

18 

8. 0 

12 

19 

2, 3 

9,10 

16 ~. 

17 

Consuiiinbles managenlent ____________________________________ 1617 

Electrical systenis anonialy -_____________________-_-__-____-- l7,18 

1\Ian;lgemellt co11trols ____________________-___--_____-__-__---- 10 

Data loas______-______-_-____----_--__-____-____~____-______- 13: 14 

JIissioii events after accident : 

Fuel cell current increase _____________________________________ 10 

B’uel cell failure __________________________________________-___ 14-13 

Fuel cell failure _________________________________________---_ 13 11 tank So. 2 telemetry ___________________________________ 11.12 

Saturn V/S-11 stage anoiiialy _________________________________ 10 

Service iiioclule heating 18 

Spacecraft niotioiis 12.13 

Taiili fan utilization __________________________________________ 10.11 

OJISF’s response to recommeiiclatioii~ 3s. 59 

Oxygen pressure restoration attempts 15.16 

Oxygen shelf assembly S 

Oxygen storage tanks : 

Structure 6. 7 

Oxygen tank So. 2: 

Design and functions Z~ 35-3s 

Failure ___-_____-______-_-__-_- - 4, .:

Preflight anomaly 10 

Preflight wiring damage G 

Telemetry 11.1‘7 

Prrssure 8 

Reaction Control System (RCS) status 1.7 

Recommeiiclatioas ___________________________________ 20. 49, 50. 59, 60. 63 

Summary__-___--_-_-_-_-------_-------_-___---___-_-________ 2. 3. 21-27 

Testing and analysis status _-_____________________________________ h2 

Text of chapter .? : “Findings, Determinations. and Recommendations“ 2-20 

Trajectorr changes---------------- IS 

Transfer to V.S.S.R 53 

hpollo 13 accident comparison 6% 70 

Fire prerention design philosoiilir ________________________ _____ 71. X2 

Subsgstein review ------------------ ____-_____ - 70 

JIission erents through accident : 

Apollo Saturn 2M : 
112 

Page 

TS 

6 F 

37 

1 .i r- 11 

43 

43 

Committees and boards : 

Appropriations Committee, House of Representatires _______________ 74 

Lunar and Planetary Missions Board ___________________________ 100, 101 

Space Science Board, T\’AS_-_--__--__-___------_-_--_--_------__-- 101 

101 

98 

101 

Conferences : American Association for AdTancement of Science, Washington, D.C., December 1969 ____________________------__-_-_-____----

Construction of facilities: Budget cnt, fiscal year 1971 ______---_____--- 

Cornell University : Future lunar missions endorsement ________________ 
113 
114 

E 

ECS. (See Environnieiital Control System (ECS).) 

Electrical arcs : Ignition ~~otential____________-___________----- 

Electrolytic batteries : Short-circuiting __________________----------- ---- 

Environmental Control System (EC8) _________________________ -------- 

Fiberglass cloth. (See Beta cloth.) 

Fill tube assembly. (See Quantity probe, oxygen storage tanks/Pill tube 

Firr prevention. ( Sec Aiiollo spacecraft/Fire prevention.) 

”Fire Prerention in hliollo” (test) 

Flight coreralls : Materials develol~iiient 

Fra Jlauro crater (moon) ________________________________________--- Fuel cells. (Set Apollo 13/Fuel cells.) 

Funding 

Fuses : Blon- tiiiie _____________________-__-_-______-__-_-____-_--__-_ 

F 

assembly. ) 

c; 

Gelirig. James J. : Coinnients : 

dpollo Saturn 204 : Fire prereiition philosophy_---__-_--------- 

Apollo spacecraft : Fire prevention ............................ Rrrch Aircraft Corn. : Tlieriiiostatic s~itches testing_----------- 

Heater assembly, oxygen storage tanks : Thermostatic sn-itches 

Oxygen storage tanks : Temperature monitoring procedures---_-- 

I’olyvmer Research Laboratory, dmes Research Center : Coiistructioii of facilities Imdget cut _________________________________ 

Aerospace Safety Advisory Panel : 

testing _________________________________________---_____-_- 80 

83 

98 

Inquiries : 

Activities review _________________________________________ 93-9.5 

Directive from NASA Administrator on Apollo 13 Review 

Board ______________________-_-_-----___-_-_------_-

93 

Members and their biographies ____________________________ X3 

Testing and analysis status _______________________________ 82 

Apollo Saturn 204 accident : Subsysteiii re&w impact_---__---_ 70 

Beech Aircraft Corp. : Thermostatic switches testing__-------_-_ 81 

Heater assembly, oxygen storage tanks : T~herinostatic switches 

testing ________________________________________-__----_-_-_ S1. 96 

Lunar missions : Future IiiiHsions justification _________________ $IS, 99 

Lunar module (13) power system anomaly : Cost or schedule 

impact ________________________________________-__-_--_---- 68 

Osygrn storage tanks : 

Designer _________________________________________---_--- 81 

Temperature monitoring procedures _______________________ 82.83 

Vnsealed fan motor elimination ___________________________ I’olymer Research Laboratory. hies Rrsearch Center : Construction justification _________________________________________-- Teflon : 1,mitioii tests ____________________-_---

Written nuestions answered br 

nr. rjnle D. llyrrc: : 

Apollo slincecraft, Aliollo 13 ReT-ieiv Board inodification rw- ommendations ____________-___________________________- Apollo spacecraft, pressure ressels review status___-___-_- 

dpollo spacecraft. review and eraluatioii of co-location of 

redundant subsystems _________________-________________ 

dpollo spncecraft, systems elenients evaluation. tests and 

analyses _________________________________________----- Serrice inodule ( SM) , components/subsystenis shock qualification testin~_-___--_----__------__--------------------- 

Geological Swiety of America : Future lunar missions endorsement-__- 

George C. Marshall Space Flight Center, Huntsrille, Ala. : Apollo &pacecraft siibeystems reriew ________________________________________--_- 
115 
116 

Holland. Senator Spessard L. : Comment : Page 

compared to KSC----__--_-----_-___-___-___________-___-__ 76 

reductions ________________________________________-______- 74. 7; 

compared to KSC -_________________________________________ 76 

Sew testing procedures-_____-__-__-___-________-___--____ 74 

Swiixtthetic damage. cause and precautions ________________ 72. 73 

Third tank spacing problems ______________________________ 73 

Teflon eoTered wires: Fire hazard _____________________________ $3. 74 

Hornbeck. Dr. John Austin : Biography 

Hmse of Representatives. (See Appropriations Committee, House of 

Reimsentatires) -_-_______________________________________-

90, 91 

JIicsion Control Center, JISC : Countdown location efficiency as 

Aiipropriations and budget : Research and derelopnient budget 

JIission Control Center. JISC : Countclon-n location efficiency as 

Oxygen storage tanks : 

Inquiries : 

I 

Inconel tube. oxygen storage tanks : Possible failure site ____________-___ 7,42 Indian Ocean_________-_______________________--_____________________ 43 

Inwlation. (See Kapton : RIylar : Oxygen storage tanks : Teflon.) 

Interagency cooperation : Aariciilture Departnient/SBSB : Lunar soil research ________________ 103 

r.S.S.R./I’.S. : Apollo 13 Reriew Board report submission ___________ 63 

J 

John F. Kennedy Space Center, XASA, Kennedy Space Center, Fla. : .kpollo spacecraft subsystems review ___________________-___________ 

Conntdown location efficiencF as compared with JfSC _______________ Oxrgen tank So. 2 : 

.58 

76 

Detanking procedures __________________________________ 2-4,6,24,41 

Ground tect operations _____________________________________ 2,46,45 

Heater operation__________________-___-_____-________-----___ 40 

Thermostatic switch testing _______________________ 4, 6, 23-26 47, 80, 81, 96 

K 

Kapton : Thermal insulation __________________________________________ KSC. (Gee John F. Kennedy Space Center, NASB, Kennecly Space Center, 42 

Fla.) 

L 

Laqers : Earth rotation measurements ............................. --___ 

Launch rehicleb. ( Sc-e Saturn V launch vehicle.) 

Lewis Recearch Center, Cleveland. Ohio, (See ,lerospace Safety Research 

“Lifeboat” niodeq. ( See Coinmand Nodule ((211) : Lunar Nodule (LM) .) 

LJI. (See Lunar Module (LJI) .) 

Low. Dr. George 31.. General testimony : 

1,unar and Planetary Jfissions Board : Future lunar missions jnstiLunar missions : 

103 

and Data Inctitute IASRDI) .) 

Beta cloth (Fiberglass cloth) : Comparison with Teflon---------__- 78 

Somex/Durette : Terms clarification ..................... ------___ 77, $8 

ficatioll ___________-____________________________-___------------- 100,101 

Future misqionq justification : Expected benefits of lunar exploration ____________________-- 103-104 

Hardware availability _____-___________________________ _______ 1@ 

Opinions of the Scientific Community ________________________ 100-103 

Prime arguments ________________________________________---- 100 

Summary ________________________________________-_-

98,99 

International interest ________________________________________-___ 103 
117 

Page 

-ictiration ...................................................... 16-13 

Attitude control system deactiratioII_______________________________ 16 

Carboll dioxide system modification _______________________________ 17 

Consumlibles management .................................. 16.17.44,37 Descent propiilsion system utilization _____________________---- ----- 43 

Electrical system anomaly ........................................ 17,lS 

In-fliglit anomaly ___________________________-_-------- ------------ -13 

“LifebogLt” mocle study ________________________________________- 62,3*5,67 

Power s,ystem anomaly ___________________________________________ GS 

103 

91 

Lunar hlodule (LM) : 

Lunar region. (See Fra Mauro crater (mooii).) 

Liinai. soil rosearch : Agriculture Department/S.iSh cooperatio~i________- 

Lundin, Brrice T. : Biography ________________________________________- 

JI 

JIanned Slxwecraft Center. Houston. Tes. (scc also dpollo 13 Inr-estigation 

ilccideni; investigation _________________________________________--- 2 

Apollo 13 Inrestigation Team activities ____________________------ -- 29 

Apollo 3 3 Rerien Board recommendations _________________________ 49, 50 

Apollo spacecraft wbsystem rerien- _______________________________ .?S 

Countdon7n location efficiency as compared vith KSC ________________ 76 

LJI flight anomaly inrestigation ___________________________________ 17,lS 

JIission Control Center i3ICC) : 

Carbon dioxide system modification ___________________________ 17 

Data handling and eraluation ________________________________ 13-15 

Data loss, Apollo 13 ________________________________________-_ 13.14 

Electrical system management ______________________--____-_-_ 1.7 

LlI activation checli list nioclificntion .......................... 16 

LJI consuinables management ................................. 16.17 

Osxgen pressure restoration attempts __________________________ 15.16 

Second ler~1 limits sensjng .................................... 57 

Warning systems reriev- and modifications ..................... 64. 6.7 1 storage tank testing ....................................... 9.10 

rl? vessel systems hazards study ............................. 10 

Sincecraft emergency equipment reyieTy ........................... 52 

Thermoqtatic sn-itch testing ______________________________________ 25 

Jlassachuseits Institute of Technologg (JIITI : Future lunar misfions 

endorsement ___________________________________________-__-___-_-- 101 

JICC. iSre JIanned Spacecraft Center. Hoiiston. Tes./Jlission C‘nntrol 

Center (JICC).) 

JIicsion Control Center. (See JI~nnrd Spacecraft Center. Hou-ton. Tes./ 

JIission Control Center (JICC).) 

VIT. iScr Jlhssachusetts Institute of Teclinolo,cj- i1IlT). 1 

JISC. i8ce JIanned Spacerraft Center. Howton. Tes.) 

JISFC. (See George C. Marsha11 Space Flight Center. Hmntsville. dln. i 

3Iyers. Dr. 1)ale I). 

Team, JIS’C : Test Specification Criteria Document (TSCD), JISC 1 : 

Biowmhy -----------------------------__--__---_-__-__-_--____-- 10s 

General te.stiniong : -- .ipclllo program : Engineering teams rerien- ..................... Apr)ropriatioiis and budget : Re.cearch and de~-elolmienr iindget 

JIission Control Cmter. JISC : Countdown location PffiPienrT 2s 

is 

reductions _-_-______________________-__~____~______-____-__ 7.5 

Review Board report recommendations ...................... 68,50 

Oxygen storage tanks : Sympathetic damage, cause and precautions __________________________________________-__________- 72, 7.1 

Teflon corered n-ires : Use discontinuance ______________________ 73, 74 

Apollo spacecraft, $pol10 13 Reriew Board modification recommendations ___________-____________________________________ 9i 

Apt)llo spacecraft, pressure ressels rerim- status _______________ 98 

Apollo spacecraft. review and eraluation of colocation of re- dundant subsystems _j ________________ 97 

n’ritten answers to questions submitted by James J. Gehrig : 
118 
119 

Pace 

1' 

General testiiiiony : 

Aerob1mce Safety Advisory Panel : RerieIy of SASA manage- 

121 

0