THE APOLLO 13 ACCIDENT 

HEARINGS 

COMMITTEE ON 

SCIENCE AND ASTRONAUTICS 

U,S, HOUSE OF REPRESENTATIVES 

IBEBORE THE 

NINETY-FIRST CONGRESS 

SECOND SESSION 

JUNE 16, 1970 

[No. 191 

Printed for the use of the 

Committee on Science and Astronautics 

U.S. GOVERNMENT PRINTING OFFICE 

47-591 0 WASHINGTON : 1970 
COJIJIITTEE OS SCIESCE ASD ASTROSAUTICS 

GEORGE P. MILLER, California, Chairmarr 

OLIS E. TEAGUE, Texas 

JOSEPH E. KARTH, Minnesota 

KES HECHLER, West Virginia 

EhIILIO Q. DADDARIO, Connecticut 

JOHX W. DAVIS, Georgia 

THOMAS N. DOWNING, Virginia 

JOE D. WAGGONSER, JR., Louisiana 

DOS FUQUA, Florida 

GEORGE E. BROWS, JR., California 

EARLE CABELL, Texas 

BERTRAM L. PODELL, New York 

WAYNE N. ASPISALL, Colorado 

ROY A. TAYLOR, Xorth Carolina 

HENRY HELSTOSKI, New Jersey 

XiRIO BIBGGI, New York 

JAMES W. SYXINGTON, Missouri 

EDWARD I. KOCH, New York 

JBJIES G. FCLTOS, Pennsylvania 

CHARLES A. NOSHER, Ohio 

RICHARD L. ROUDEBUSH, Indiana 

ALPHOKZO BELL, California 

THOJfAS 11. PELLY, Washington 

JOHN W. WYDLER, Sew Tork 

GUT T'ASDER JAGT, Michigan 

LARRY WIKN, JR., Kansas 

D. E. (BUZ) LUKENS, Ohio 

ROBERT PRICE, Texas 

LOWELL P. WEICKER, JR., Connecticut 

LOUIS FREY, JR., Florida 

BARRY >I, GOLDWATER, JR., California 

CHARLES F. DVCASDER, Czecictive Director and Chief Counsel 

JOHS A. C.ARSTARPHES, Jr., Cltief Clerk and COitnsel 

PHILIP B. TEAGER, Counsel 

FRAsK R. HAMXILL, Jr., Coztnsel 

11'. H. BoosE, Tecliiiical Consultant 

JAMES E. WILSOS, Tecltnical Consultant 

RICHARD P. HIKES, Staff Consultant 

HAROLD A. COULD, Technical Consztltant 

PHILIP P. DICKISSOS, Technical Consultant 

W~LLIAM G. WELLS, Jr., Technical Consultant 

K. GUILD SICHOLS, Jr., Staff Consultant 

ELIZABETH S. KERSAS, Scientific Research Assistant 

FRAKK J. GIROUX, Clerk 

DESIS C. QUIGLET, Publicatioits Clerk 

RICHARD K. SHULLAW, Assistant Publications Clerk 

J. THOXAS RATCHFORD, Science COllStLltant 

JaJfEs A. ROSE, Jr., Minoritg Staff 

(11) 
CONTENTS 

STATEMENTS 

June 16, 1970: Page 

Dr. Thomas 0. Paine, Administrator, NASA _________..__2 

Edgar M. Cortright, chairman, Apollo 13 Review Board [Director, 

Langley Research Center) ; accompanied by the following members 

of the Apollo 13 Review Board: 

Robert F. Allnutt (assistant to the Administrator, NASA HeadXeil Armstrong [Astronaut, Manned Spacecraft Center) 

Dr. John F. Clark (Director, Goddard Space Flight Center) Brig. Gen. Walter R. Hedrick, Jr. (Director of Space, DCS/ 

Tincent L. Johnson (Deputy Associate Administrator-EngineerMilton Klein (manager, AEC-NASA Space Nuclear Propulsion 

George Malley, counsel, Apollo 13 Review Board (chief counsel, 

quarters) 

R. & D., Headquarters, USAF) 

ing, Office of Space Science and Applications) 

Office) 

Langley Research Center) ______________.._

3 

PREPARED STATEMEST 

Edgar 11. Cortright, chairman, Apollo 13 Review Board [Director, Langley 

Research Center) ______.____________.__

3 

APPEXDIX 

Summary volume of Apollo 13 Review Board Report _______.__._

83 

(111) 
THE APOLLO 13 ACCIDENT 

TUESDAY, JUNE 16, 1970 

HOUSE OFREPRESENTATIVES, 

COMMITTEE ON SCIENCE AND ASTRONAUTICS, 

Washington, D.C. 

The committee met, pursuant to notice, at 10 :05 a.m., in room 2318, 

Rayburn House Office Building, Ron. George P. Miller, chairman, 

presiding. 

The CHAIRMAN. The committee will be in order. 

Dr. Paine, hfr. Cortright, members of the Apollo 13 Review Board, 

we are pleased to welcome you to the committee today for the purpose 

of presenting findings, determinations, and recommendations of the 

A 0110 13 Review Board. 

%r. Paine, I would like to commend you for the appointment of a 

most competent and outstanding board to review the Apollo 13 accident and the circumstances surrounding it. 

Mr. Cortright, whom this committee knows well, has distinguished 

himself not only as an administrator in the NASA Headquarters organization but also as a field center director at Langley Research 

Center. 

The other members of the board are similarly well qualified to have 

participated in this intensive and searching review. 

As I stated at the time of the Apollo 13 accident the committee decided that sufficient time should be allowed for NASA to fully investigate the accident and at such time that this investi ation was comthe accident. 

Therefore, I have asked you to appear here today even though the 

board’s report was only submitted to you yesterday, Dr. Paine, because 

I feel it is important that the members of the committee receive a firsthand and timely review of the Apollo 13 accident. 

Dr. Paine, I understand you have a short statement and then Mr. 

Cortright will go into the details of the accident and the board’s 

findings. I want to give all members an opportunity to ask questions, so will 

you please proceed. 

Before proceeding, I would also like to make a part of the record 

the fact that Mr. Wilson of the staff of the committee was appointed 

to act as an obsen-er with the boa.rd. 

I want to thank you for the courtesies you have shown us and it has 

given us a new system of liaison. Please proceed, Dr. Paine. 

pleted the committee would convene to receive NAS a ’s evaluation of 

(1) 
2 

STATEMENT OF DR. THOMAS 0. PAINE, ADMINISTRATOR, NASA 

Dr. PAINE. Mr. Chairman and members of the committee, on April 

17, Dr. George Low and I established the Apollo 18 Review Board 

under the direction of Mr. Edgar &I. Cortright, director of the Langley Research Center. The instructions to the board are contained in a 

memorandum dated April 17, and the membership of the board in a 

memorandum dated April 20, 1970. ~vhich are reproduced in the summary volume of the report you have received. 

The past 2 months have involved long hours and very hard work by 

the review board and supporting elements in NL4SA and the industrial community. I would like to take this opportunity to extend my 

thanks to them for the thoroughness of their investigation and their 

dedication to this arduous assignment. 

Since I received the review board report only yesterday, I hare not 

had a chance to review it in detail. Nor have I had the benefit of the 

independent assessment which is being carried out by the Aerospace 

Safety Advisory Panel, chaired by Dr. Charles Harrington. 

The Office of Manned Space Flight is also conducting a separate 

review of the report. 

In about 10 days I will receive the results of the safety panel and the 

manned space flight review. Until I have received and studied these reports. I will obviously not be in a. position to give you my evaluation 

of the board‘s recommendations or NASA‘s future actions. 

Earlier re announced a change in our lunar landing schedule involving a delay of the Bpollo 14 launch from October to the December 

launch window. However, this is subject to rei-iexT in light of the report of the Apollo 13 Reiyiew Board and we n-ill not fly Apollo 14 to the 

moon until we are confident that we have done everything necessary 

to eliminate the conditions that caused or contributed to the problems 

on Apollo 13. 

I believe that, as we plan man’s future course in space, the preface 

to this report should be a reminder of the nature of the challenge rre 

have undertaken. Let me quote : 

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 vehicle, 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 success- ful lunar surface explorations is a tribute to those men and women who con- ceived, designed, built, and flew it. 

Perfection is not only difficult to achieve, but difficult to maintain. The imperfection in hpollo 13 constituted a near disaster. adverted only by outstanding performance on the part of the crew and the ground control team ivhich supported them. 

The Board feels that the nature of the Apollo 13 equipment failure ho!ds important lessons rrhich, TThen applied to future missions, will contribute to the 

safety and effectiveness of manned space flight. 

Afr, Chairman, there has been time for me to reach one conclusion on 

the report of the Apollo 13 Review Board, and that is that the board 

and their supporting teams have done a. magnificent piece of technical 

detective work that carefully reconstructs the background and the 

events xhich took place aboard Apollo 13 800,000 miles from earth. 
3 

1 JT-ould 110~ like to introduce the chairman of the rerieJT board, 

Mr. Edgar Cortright, who will briefly discuss the report and respond 

to your questions. 

The CH~IIRMAX. Thank you T-ery much, Dr. Paine. 

FFTe are very happy to have you here, Mr. Cortright. 

STATEMENT OF EDGAR M. CORTRIGHT, CHAIRMAN, APOLLO 13 

REVIEW BOARD ; DIRECTOR, LANGLEY RESEARCH CENTER 

Mr. CORTRIGHT. Thank you, Mr, Chairman. 

I have a prepared statement and, vith your permission, I will submit this for the record and attempt to convey to sou JThat the Board 

has done and what our conclusions hare been in a more informal 

manner. 

The CHAIRMAS. Without objection, that rill be the manner in which 

m-e will proceed. 

(The prepared statement of Mr. Cortright is as follows :) 

PREPARED STATEMEST OF EDGAR M. CORTRIGHT, CHAIRJIKX, 

APOLLO 13, REVIEW BOARD, KATIOSAL AEROSAUTICS AISD SPACE 

ADJIISISTRATION 

Air. Chairman and Members of the Committee : I appreciate this opportunity to appear before the Committee to summarize 

tlie Report of the Apollo 13 Review Board. 

As you know, yesterday I presented this Report on behalf of the Board to 

the Administrator and Deputy Administrator. Copies of the Report vere given 

to the Members and Staff of the Committee, and the Report was made public yesterday afternoon, at which time Dr. Paine and I held a press conference. 

This morning I would like first to outline for the Committee how the Board was established and how 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 events, the 

events of the accident itself, and the recovery procedures rhich were implemented to return the crew safely to earth. I will 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. I will report these to you and be pleased to answer any 

questions you may have on the Board’s work. 

ESTABLISHXEST AXD HISTORY OF THE BOARD 

The Ago110 13 Review Board was established, and I was appointed Chairman, on April 17. 1970. The charter of the Board was set forth in the memorandum 

ivhich established it. Under this charter the Board was directed to : 

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

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

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

determines to be significant and relevant, including studies, findings. recom- mendations, and other actions that have been or may be undertaken by tlie program offices. field centers, and contractors involved. 

(c) Direct such further specific inrestigations 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 fiight and ground recovery actions. 
4 

(e) Develop recommendations for corrective or other actions, based upon its 

(f) Document its findings, determinations, and recommendations and submit 

The Membership of the Board mas established on April 21, 1970. The members 

findings and determinations or conclusions derived therefrom. 

a final report.” 

are : 

Mr. Edgar M. Cortright, Chairman (Director, Langley Research Center) 

Mr. Robert F. Allnutt (Assistant to the Administrator, NASA Hqs) 

Mr. Neil Armstrong (Astronaut, Manned Spacecraft Center) 

Dr. John F. Clark (Director, Goddard Space Flight Center) Brig. General Walter R. Hedrick, Jr. (Director of Space, DCS/RCD, Hqs. USAF) 

Jlr. Vincent L. Johnson (Deputy Associate Administrator-Engineering, Office of 

>I?. Milton Klein (Manager, AEC-SASA Space Nuclear Propulsion Office) 

Dr. Hans 11. Jlar,k (Director, Ames Research Center) 

Legal Counsel to the Board is Mr. George T. Nalley, Chief Counsel, Langley Research Center. 

Appointed as Observers were : 

Xr, William A. Bnders (Executive Secretary, Sational Aeronautics and Space 

Council) 

Dr. Charles D. Harringtan (Chairman, SASA Aerospace Safety Adrisory Panel) 

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

Lewis Research Center) 

Ur. James E. Wilson, Jr. (Technical Consultant, House of Representatives, Committee on Science and Astronautics) 

The documents establishing the Board and its membership and other relevant 

documents are included in Chapter 1 of the Board’s Report. The Review Board convened at the Manned Spacecraft Center (AISC), Houston, Texas, on Tuesday, April 21, 1970. Four Panels of the Board mere formed, 

each under the overview of a member of the Board. Each of the Panels mas chaired by a senior official experienced in the area of review assigned to the 

Panel, In addition, each Panel mas manned by a number of experienced spe- cialists to provide in-depth technical competence for the review actirity. During the period of the Board’s activities. the Chairmen of the four Panels v-ere respon- sible 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 as- suped specific responsibilities related to the overall review. 

Space Science and Applications) 
APOLLO 13 

REVIEW BOARD 

I E. M. CORTRIGHT. CHAIRMAN 

I I I 

COUNSEL SECREIARIAI 

C. W. MAIHEWS G. 7. MALLEI E. P. SWIEDA 

OMSF IECHNICAL SUPPORI 

R. F. AUNUll 

N. A. ARMSTRONG M. KLEN 

J. F. CLARK 

W. R. HEDRKK 

V. L JO

H. M. M

c. D. I. HARRINGTON I. PINKEL 

1. E. WILSON, JR. 

1. G. 

E

I I L 

I 

II 

I 1 I MANUFACIURING 

h IESI MISSION EVENIS 

I. 

L J 

I 1. 1. WILLIAMS I 

I I. B. BALLARD I 

EVENIS 

1 M. P. FRANK 1 

E. F. BAEHR 

SUBSYSIEM h 

SYSIEM IESIING 

K. 1. HEIMBURG 

b 8. I. MORRIS 

RELlABlLlIV h 

OUALIN ASSURANCE 

1
6 

Team Work and the Apollo 13 Review Board activibies was effected through 

the Manned Space Flight Technical Support official and by maintaining a close 

and continuing working relationship between the Panel Chairmen and officials of 

the MSC Investigation Team. In addition, Board members regularly attended daily status meetings of the Manned 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 

analysis 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 NSC, but in addition, the Board established an extensive series of special tests and analyses at other NASA Centers and at 

contractor facilities. Members 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 which 

inrolred several hundred people at its peak, the elements of the inflight accident were reproduced. All indications are that electrically initiated combustion of 

Teflon insulation in oxygen tank So. 2 in the service module vas the cause of 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 combustion test, demonstrated the most probably tank failure mode. 

Simulated tank rupture tests in a 1/2 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 have been generated. I hare with me a brief film, highlighting these tests, which I would like to show at the conclusion 

of my statement. APOLLO 13 SYSTEJIS 

Before tracing the analyses which lead to the Board’s conclusions-and to place 

them in proper context-I would like to explain the design and functions of the 

oxygen tank #Z as a part of the Apollo system. Details of the entire ApollO/ Saturn Space Vehicle are set forth in the Report and its Appendices. 
Launch escape systeiin 

Boost protective cover 

Coilinland iiiooule 

Service niodule 

Spacecraft LM adapter 

Lunar module 

Spacecraft Space vehicle 

Vi trtir- 3- I. - Apvllo/Stc
8 

ENVIRONYENTAL 

CONTROL SUBSYSTE OLZLE EXTENSION RADIATOR 

FORWARD BULKHEAD Ills1 

OXYGENTANKS 

HYDROGEN TANKS 

WAND HIGH GAIN ANTENNA 

SERVICE PROWLSION ENGINE 

SERVICE PROPULSION SUBSYSTEM 

OXYGEN TANKS HYDROGEN TANKS FUEL CELLS 

SERVICE PROPULSION SUBSYSTEM } FUELTANKS 

CENTER SECTION, SERVICE PROPULSION ENGINE AND 

HELIUM TANKS 

Figure 3-6.- Service module. 

Slide 3 

Slide 2 shows the hpollo/Saturn Space Vehicle, with which you are all 

familiar. Slide 3 shows the service module which, as you know, is designed 

to provide the main spacecraft propulsion and maneuvering capability during a mission. It also contains most of the spacecraft consumables (oxygen, water, 

propellant, and hydrogen) and supplies electrical power. The service module 

is divided into six sectors or bays surrounding a center section. The oxygen 

tank, to which I referred, is located in Bay 4 (shown in more detail on Slide 4), 
9 

NASA -S -70-512-V 

ARRANGEMENT OF FUEL CELLS AND CRYOGENIC 

SYSTEMS IN BAY 4 

h-2 Slide 4 

along with another oxygen tank, two hydrogen tanks, three fuel cells and interconnecting lines, and measuring and control equipment. The tanks supply oxygen to the environmental control system (ECS) for 

the astronauts to breathe, and oxygen and hydrogen to the fuel cells. The fuel 

cells generate the electrical power for the command and service niodules during a mission. The next slides (Sos. 6, 6 and 7) are photographs of Bay 4 of the 

service module for Apollo 13, showing the major elements and their interconnection. Slide 7 shows the oxygen tank #2 in place. 
10 

Slid. 5 
11 

6 
12 

1 As the simplified drawing in Slide S indiaates, each oxygen tank has an outer 

shell and an inner shell, arranged to provide a vacuum space to reduce heat 

leak, and a dome enclosing paths into the tank for transmission of fluids, and 

electrical power and signals. 
13 

Blowout disc Closeout cap 

Supply 

line - 

To fuel 

cell/ ECS 1 

Pressure 

0 

transducei 

Figure 4-1.- Oxygen tank no. 2 internal components. 

slide 8 

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 thermostatically protected heater coils and tn.0 small 

fans driven by 1800 RPJI 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 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 temperature 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. 

47-591 0-70-2 
14 

n 

Temperature sensor 

Quantity probe. 

-Thermostat 

Heater 

Figure 4-2.- Oxygen tank wiring and lines. 

Slide 9 

The oxygen tank, as designed, contained materials, which if ignited will burn 

in supercritical oxygen. These include Teflon, used, for example, to insulate the 

miring, and aluminum. 

Pressure in the tank is measured by a pressure gage in the supply line, and a pressure sn*itch 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. 

The oxygen tank is designed for a capacity of 320 pounds of supercritical 

oxygen at pressures ranging between 565 and 933 pounds per square inch absolute (psia). The tank is initially filled with liquid oxygen at -297’F and 
15 

operates over the range from -340” to +SOOF. The term “supercritical” means 

that the oxygen is maintained at a temperature and pressure which assures that it is a homogeneous, single-phase fiuid. 

The burst pressure of the oxygen tank is about 2200 psia at -160”F, over twice the normal operating pressure at that temperature. 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 pressure 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. 

As shown in Slide 9, each heater coil is protected with a thernioatatic switch, 

mounted on the heater tube, which is intended to open the heater circuit when 

it senses a temperature of 80” F. As I ?\-ill point out later in tracing the Board’b 

concludons as to the cause of the accident, when the heaters were powered from a 65 volt DC supply at KSC during an improvised detanlring procedure, these 

thermostatic switches, because they were rated at only 30 T’ DC, could not 

prevent an overheating condition of the heaters and the associated wiring. Tests 

conducted for the Board indicate that the heater tube assembly was probably 

heated to a temperature of as much as 1000° F during this detanking procedure. 

THE APOLLO 13 MISSIOR’ 

With this general background, I mill now summarize the Apollo 13 mission. 

This mission, as you kno\v, was designed to perform the third manned lunar 

landing. The selected site was in the hilly uplands of the Fra JIauro formation. A 

package of fire scientific experiments was planned for emplacement on the 

lunar surface near the lunar module landing point. Additionally the dpollo 13 

landing crew was to gather the third set of selenological sample3 of the lunar 

surface for return to earth for extensive scientific mialysis. Candidate future 

landing sites were scheduled to be photographed from lunar orbit. The crew consisted of Captain James A. Lorell, Commander, Fred 77’. Haise. Lunar 

Module Pilot ; and John L. Sn-igert, Jr., Comniand Jlodnle Pilot, n ho replaced 

Thomas K. Jlattingly, 111, who had been espoard to rubella and, after tests, 

found not to be immune. 

Launch was on time at 2 :13 p.m., EST on April 11 from the KSC Launch Complex 396. The spacecraft was inserted into a 100-nautical mile circular earth 

orbit. The only significant launch phase anomaly \vas premature shutdown of 

the center engine of the S-I1 second stage. This anomaly, although serious, mas 

not related to the subsequent accident. It is being investigated by the Apollo or- ganization. As a result of this shutdown, the remaining four S-I1 engines burned 

34 seconds longer than planned and the S-IT’B third stage engine burned a ferr 

seconds longer than planned. At orbital insertion, the \-elocity was within 1.2 

feet per second of the planned velocity. Moreover, an adequate propellant mar- gin was maintained in the S-IT’B for the translunar injection burn. 

After spacecraft systems checkout in earth orbit, the S-IVB restarted for 

the translunar injection (TLI) burn, with shutdown coming some six minutes 

later. After TLI, Apollo 13 vas 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-IT’B 

about three hours into the mission, and after a brief period of station-keeping. 

the crew maneuvered the CSX into dock with the LU vehicle in the L3I adapter 

atop the S-IVB stage. The S-IT’B stage Tas separated from the docked CSJI and 

LJI shortly after four hours into the mission, and placed on a trajectory to ultimately imwct the moon near the site of the seismometer emplacecl by the Apollo 

12 crew. 

At 30 :40 :49 g.e.t. (ground elapsed time) a midcourse correction nianeuver 

 as made using the service module propulsion qysteni. This maneuver took 

Apollo 13 off a free-return trajectory and placed it on a non-free return trajectory. A similar profile had been flown on Apollo 12. The objective of leaving a free-return trajectory is to control the arrival tirile at the moon to insure the 

proper lighting conditions at the landing site. The transfer maneuver 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 ivas normal, so far as telenietered data and crew ob-ervations 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 
16 

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 :% 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 

g.e.t. Astronauts Love11 and Haise were cleared to enter the LJI to commence inflight inspection for the LlI. Sfter this inspection period, the lunar module was powered down and preparations were underway to close the LN hatch and run through the presleep checklist when the accident in oxygen tank =2 

occurred, 

At about 66:63, flight controllers in the Mission Control Center at JISC re- quested the crew to turn on the cryogenic system fans and heaters, since a 

master alarm on the CJI Caution and Warning System had indicated a low 

pressure condition in the cryogenic hydrogen tank #1. This tank had reached 

the low end of its normal operating pressure range several times previously 

during the flight. S-cigert acknowledged the fan cycle request and data indicate 

that current was applied to the oxygen tank #2 fan motors at 33 :53 :20. 

About 2% minutes later, at 65 :54 :63.5, telemetry from the spacecraft was lost 

almost totally for 18 seconds. During the period of data loss, the Caution 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 if1 or its lines or valves. Photographs of the serrice module taken by the crew later 

in the mission (Slide 10) show the panel missing, the fuel cells on the shelf above 

the oxygen shelf tilted, and the high gain antenna damaged. 
17 

The resultant loss of oxygen made the fuel cells inoperatire, leardng the CAl 

Iyith batteries iiornially used only during reentry as the sole power source and 

with only that oxygen contained in a surge tank and repressurization packages. 

Tlie lunar module, therefore, becanie the only source of sufficient battery pwer 

and oxygen to permit safe return of the crew to earth. 

SCJIl\IARY ASALYSIS OF THE ACCIDEST 

Tlie Board determined that conibustion in oxygen tank #2 led to failure of that 

tank, damage to oxygen tank #1 or its line3 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 niissioii. 111 tlie attempt to determine 

the cause of ignition in oxygen tank #2, the course of propagation of the combustion, the inode of tank failure, and the may in which subsequent damage 

occurred, the Board lias 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 provide x clear picture of the nature of the accident and the events 

which led up to it It is now apparent that the extencled heater operation at 

KSC: damaged the insulation on \Tiring in the tank and that this set the stage 

for the electrical lsliort circuits vliich iiiitiatecl conibustion within the tank. 

While the exact point of initiation of combustion and the specific propagation 

path involved niay never be 1woir.n with certainty, the nature of the occurrence is sufficiently )Tell unclerstoocl to permit taking corrective steps to prevent its recurrence. The Board lias identified the most probable failure mode. 

The following discussion treats the accident in its key phases : initiation, propagation and energy re1ea.e. loss of oxygen tank #2 system integrity, and loss of 

oxygeii tank #1 systeni integrity. Slide 11 show the key events in tlie sequence. 
18 

v) 

ti? 

ILu 

Z 

@ 

V 

V 

Q 

Initiation 

The evidence 

, 

points strongly to an electrical short circuit with arcing as the initiating event. Xear the end of the 55th hour of flight, about 2.7 seconds 

after the fans were turned on in the SX 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 decreased by an amount consistent with the loss of power to one fan. So 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 quantity gage and temperature sensor are very low power devices. The next anomalous event recorded was the beginning of a pressure rise in oxygen tank #2, 13 seconds later. Such a time lag 

is possible with lorn level combustion at the time. These facts point to the likelihood that an electrical short circuit with arcing occurred in the fan motor on its leads to initiate the accident sequence. The energy available from the short 

circuit is estimated to have 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. 

n 
19 

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 launch. The likelihood of damage and the possibility of electrical ignition have been verified by tests. 

Propagation 

TVhile 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 

pomer to account for all of the energy requird to produce the observed pres- sure 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 reactive 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 would 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 about 13 seconds after the 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 slow 

propagation of burning along the Tviring. it is likely that the region immediately 

around the temperature sensor did not become heated until this time. 

The data on the combustton of Teflon in supercritical oxygen in zero gravity, developed in special tests in support of the Board, indicate that the rate of 

combustion is generally consistent with these observations. 

Loss of Oxygen Tnnk #2 System InttgritU 

After the relatively slow propagation process described abore took place, there was a relatively abrupt loss of oxxgen tank #2 integrity. About GO seconds after 

the pressure began to rise, it reached the lieab recorded, 1008 psia, the pressure 

at which the cryogenic oxygen tank relief valve is designed to be fully open. 

Pressure began a decrease for 8 seconds, dropping to 996 psia before readings were lost. About 1.83 seconds after the last presumably valid reading from 

Tithin the tank (a temperature reading) and .8 seconds after the last presum- ably valid pressure reading (which may or may not refiect 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 approximately .42 seconds after the last 

pressure reading and approximately 38 seconds before the loss of signal. These facts all point to a relatively sudden loss of integrity. At about this time, several 

solenoid ralves, 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 was reacquired, all instrument indicators 

from osygen tank #2 were off-scale, high or low. Teniperatnres recorded by 

sensors in sereral different locations in the serl-ice module showed slight increases in the several seconds following reacquisition of signal. 

Data are not adequate to determine precisely the way in which the oxygen 

tank #2 system failed. Howerer, available inforniation, 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 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 \Tires 

enter the pressure vessel. It is likely that the combustion progressed along the wire insulation and reached this location 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 the tank dome, which is equipped with a rupture disc rated at 76 psi. Rupture of this disc or of the entire dome would then release oxygen, accom- panied by combustion products, into Bay 4. The accelerations recorded were probably caused by this release. 
20 

Release of the oxygen then began to rapidly pressurize the oxygen shelf 

space of Bay 4, If the hole formed in the pressure vessel mere large enough and formed rapidly enough, the escaping oxygen alone would be adequate to 

blow off the Ray 4 panel. However, it is also quite possible that the escape of 

oxygen was accompanied by combustion of Mylar and Kapton (used extensively as thermal insulation in the oxygen shelf compartment and in the tank dome) 

which would augment the pressure caused by the oxygen itself. The slight telmperature increases recorded at various locations in the service iiiodule indicate that 

combustion external to the tank probably took place. The ejected Bay 4 panel then struck the high gain antenna, disrupting communications from the spacecraft for the 1.8 seconds. 

Loss of Oxygen 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 mas regained, data 

show that the tank #1 system had lost its integrity. Pressure decreases mere 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 -i-alre or relief Palre. 

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 associated with the failure of the oxygen tank #2 system 

could have ruptured a line to osygen tank #l, or have caused a valve to leak 

because of mechanical shock. 

APOLLO 13 RECOVERY 

Understanding the Problem 

In the period immediately following the Caution and Warning Alarm for 

Main Bus B undervoltage, and the associated “bang” reported by the crew, the cause of the difficulty and the degree of its seriousness v-ere not apparent. 

The 1.8-second loss of telemeterecl data accompanied by the switching of the 

CSM high gain antenna mounted on the SU adjacent to Bay 4 from narrow beam width to wide beam width. The high gain antenna (HGA) does this 

automatically 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 period after data loss. In all probability, these thrusters were being fired to overcome the effects that oxygen venting and panel blorr-off 

were having on spacecraft attitude, but it was believ-ed for a time that perhaps the thrusters were malfunctioning. The failure of oxygen tank #2 and consequent removal of the Bay 4 panel 

produced a shock which closed valves in the oxygen supply lines to fuel cells 1 

and 5. These fuel cells ceased to provide power in about three minutes, when the 

supply of oxygen between the closed valves and the cells was depleted. 

The crew was not alerted to closure of the oxygen feed valves to fuel cells 

1 and 3 because the valve position indicators in the CAT were arranged to give 

warning 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 low pressure 

warning had blocked the cryogenic subsystem portion of the Caution and Warning System several minutes before the accident. A limit sense light presumably came on in Mission Control during the brief period of tank overpressure, but was 

not noticed. 

When the crew heard the “bang” and got the master alarm for low DC Main 

Bus B voltage, Love11 was in the lower equipment bay of the command module, 

stowing a television camera which had just been in use. Haise was in the tunnel between the CSM and the LM, returning to the CSII. Swigert was in the left 

hand couch, monitoring spacecraft performance. Because of the master alarm 

indicating low voltage, Swigert moved across to the right hand couch where 

CShI voltages can be observed. He reported that voltages were “looking good” at 

55 : 56 : 10. At this time, voltage on Main Bus B had returned to normal levels 
21 

and fuel cells 1 and 3 did not fail for another 1% to 2 minutes. He also reported fluctuations in the oxygen tank #2 quantity, followed by a return to the offscale high position. 

When fuel cells 1 and 3 electrical output readings went to zero, the ground 

controllers could not be certain that the cells had not somehow been disconnected 

from their respective buses and were iiot otherwise all right. 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 was realized that fuel cells 1 and 3 byere not functioning, the crew was directed 

to perform an emergency power-down to reduce the load 011 the remaining fuel cell. Observing the rapid decay in oxygen tank #1 pressure, controllers 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 

the situation became clear. 

During the succeeding period, efforts were made to Save the remaining oxygen 

in the oxygen tank #1. Several attempts were made, but had no effect. The pres- sure continued to decrease. 

It was obvious by about one-and-one-half hours after the accident that the 

osygen tank #1 leak could not be stopped and that it would won become neees- sary to use the LJ1 as a “lifeboat” for the remainder of the mission. 

By 58 :40, the LJI had been activated, the inertial guidance reference transferred from the CSM guidance system to the LJI guidance system, and the CSM 

systems were turned off. 

Return to Earth 

The remainder of the mission was characterized by two main nctirities-plan- ning and conducting the necessary proppulsion maneuvers to return the spacecraft to earth, and managing the use of consimables in such a way that the 

LM, which is designed for a basic mission with two crewmaii for n 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 07 hours 14 minutes into the mission, Haise reported hearing a “thump” 

and observing venting from the LJI. Subsequent data review shows that the LM 

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 Apollo organization is continuing. 

A number of propulsion options were developed and considered. It was neces- sary to return the spacecraft to a freereturn trajectory and to make any required midcourse corrections. Sormally, the Service Propulsion System (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 LJf descent engine if possible. The minimum practical return time was 133 hours to the Atlantic Ocean. ad 

the maximum mas 132 hours to the Indian Ocean. Recovery forces were deploy4 

in the Pacific. The return path selected was for splashdown in the Pacific Ocean 

at 112 :40 g.e.t. This required a minimum of two burns of the LJI descent engine. 

A third burn was subsequently made to correct the normal maneuver execution 

variations in the first two burns. One small velocity adjustment was also made 

n9th reaction control system thrusters. A11 burns were satisfactory. Slides 12 

and 13 depict the flight plan folloved from the time of the accident to splashd0WR. 
/--- 

K 

for entry for 

corridor cor

( 137 :4 0) (10

Figure 4-14 .- Translunar trajec
23 

m 

rl 

0) 

c 

a 

3 
24 

The most critical consumables were water, used to cool the CShI and LM 

systems during use; CSJI and LM battery power, the CSM batteries being for 

use during reentry and the LA1 batteries being needed for the rest of the mission ; 

LJI oxygen for breathing ; and lithium hydroxide (LiOH) filter cannisters used 

to remove carbon dioxide from the spacecraft cabin atmosphere. These consum- ables, and in particular the water and LiOH cannisters, appeared to be extremely 

marginal in quantity shortly after the accident, but once the LJI was powered down to conserve electric power and to generate less heat and thus use less water, 

the situation greatly improved. Engineers at MSC developed a method which 

allowed the crew to use materials onboard to fashion a device allodng the use of the CJl LiOH cannisters in the LJI cabin atmosphere cleaning system. At splashdown time, many hours of each consumable remained ai-ailable. 

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 mas 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 oxwen tank. its design. 

manufacturing, test, handling, checkout, use,-failure mode,- and eventual on the rest of the spacecraft. 

effezts 

OXYQEN TANK #2 HISTORY 

On February 26, 1966, the Xorth American Aviation Corporation, now, Xorth 

American Rockwell (SR), prime contractor for the Apollo command and serv- ice modules (CSJI), awarded a subcontract to the Beech Aircraft Corporation 

(Beech) to design, develop, fabricate, assemble, test, and deliver the Block I1 

.\pollo cryogenic gas storage subsystem. This was a follow-on to an earlier subcontract under which the somewhat different Block I subsystem was procured. 

Manufacture 

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 detected 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 flaws 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 was 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 SR on Nay 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 Xay 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 ShI 106. 

Between August 3 and August 8, 1968, testing of the shelf in the Shl was con- ducted, 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 was introduced and a decision was made to replace the complete oxygen shelf in SM 

106. An oxygen shelf with approved modifications was prepared for installation 

in SM 106. On October 21, 1968, the oxygen shelf was removed from SM 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 \vas beliered that the oxygen shelf had simply dropped back into place, and an analysis mas performed to calculate the forces resulting from a drop of two 
25 

inches. It now seems likely that the shelf was first accelerated upward and 

then dropped. The remaining bolt mas then removed, the incident recorded, and the oxygen 

ahelf was removed without further difficulty. Following removal, thie oxygen shelf was retested to check shelf integrity, including proof pressure tests, leak tests, and fan and heater operation. Visual 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 snd tests conducted during this investigation have indicated that the 

forces experienced by the shelf were probably close to those originally calcu- lated, assuming 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 could have been displaced by the event. 

The shielf passed these tests and was installed in SJI 109, the Apollo 13 service 

module, on Sovember 22, 1968. The shelf tests accomplished earlier in ShI 106 

were repeated in ,311 109 in late December and early January, with no significant 

problems, and SM 109 was shipped to KSC in June of 1969 for further testing, assembly on the launch vehicle, and launch. 

Testing at KSC 

At the Kennedy Space Center the CM and the SM were mated, checked, assem- bled on the Saturn V launch vehicle, and the total vehicle was moved to the launch 

The Countdown Demonstration Test (CDDT) began on March 16, 1970. Up to 

this point, nothing unusual about oxygen tank #2 had been noted during the 

extensive &sting ait KSC. Cryogenic oxygen loading and tank pressurization to 

331 psi was completed without abnormalities. At the time during CDDT when the 

oxygen tanks are normally vented 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, MSC, 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 66 percent. Further discussions 

between KSC, MSC, SR, 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 sleeves and tube. Such a leak would allow the gaseous oxygen 

being supplied to the vent 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 was 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 mere energized 

with the 65 volt DC GSE power supply and, about 1y2 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 pres- sure cycling technique. With the heaters and fans still energized, the tank was 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 JISC 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 between the capacitance plates of the quantity gage would 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 would have taken at least 45 hours. 

In addiition, shelf replacement would haTe had the potential of damaging or degrading other elements of the service module in the course of replacement activity.Therefore, the decision was made to test the ability to fl11 oxygen tank #2 

Wd. 
26 

on JIarch 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. 

Elow tests were first made with gaseous oxygen on oxygen tank #2 and 011 

oxygen tank if1 for comgarison. So problems were encountered, and the flow 

rates in the t\vo 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 low energy 

levels ~ould result. Then, oxygen tanks #1 and #2 were filled with LOX to 

about 20 percent of capacity on Ilarch 30 with no difficulty. Tank #1 emptied in the normal manner, but empyting oxygen tank #2 again required pressure 

cycling Xrith the heaters turned on. As the launch date approached, the oxygen 

tank #a detanking problem \ras considered by the dpollo organization. -kt 

this point, the “shelf drop” incident on October 21, 1965, at SR was iiot con- sidered and it mas felt that the apparently normal detanking which had oc- curred in 1967 at Beech was not pertinent because it was believed that a different 

procedure was used by Beech. In fact, homerer, the last portion of the procedure was quite similar, although at a slightly lower pressure. 

Throughout these considerations, which inrolred technical and management 

personnel of KSC, JISC, SR, Beech, and SASd Headquarters, emphasis was directed toward the possibility and consequence of a loose fill tube; very little 

attention was paid to the extended heater and fan operation, except to note 

that they operated during and after the detanking sequences. 

JIany of the principals in the discussions 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 escessive heat within the tank, 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 sn-itch, 

mounted on the heater tube, which is intended to open the heater circuit Then 

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 65 volt DC supply similar to the power used at KSC during 

the detanking sequence. Subsequent investigations hare shown that the thermostatic switches nsed, while rated as satisfactory for the 28 volt DC spacecraft 

power supply, could not open properly at 65 rolts DC with 6-7 amps of current. 

B rerien- of the roltage 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. Slide 14 shows a thermostatic switch melded closed after 

application of 1% amperes of 65 volts DC. 
27 

Further tests have shown that the temperatures on the heater tube subsequent 

to the switch failures may hare 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. Slides 15 and 16 show the condlition of 

wires, such as those used in the fan motor circuit, after they have been subjected 

to temperatures of about 1000" F. 

29 

Xone of the above, however, was known at the time and, after extensive con- sideration mas 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 SM and to proceed with preparations for the launch of Apollo 13. In fact, following the special de- tanking, the oxygen tank #2 was in hazardous condition whenever it contained 

oxygen and mas electrically energized. This condition caused the Apollo 13 

accident, which mas 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 assure the success of future spaceflight missions. Sereral recommendations of this 

nature are included. 

47-591 0-70-3 
30 

RECOMMEXDATIOSS 

Before reading the Board’s recommendations. I would like to point out that 

each Member of the Board concurs in each finding, determination. and reconimendation. 

The Board’s recommendations are as follows : 

1. The cryogenic oxygen storage system in the service module shonlcl be 

modified to : 

a. Remore 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 inducecl fire in the tanlr. 

b. Minimize the use of Teflon, aluminum, and other relativel;; conibnstible 

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

2. The modified cryogenic oxygen storage system should be subjected to R 

rigorous requalification program, including careful attention to potential operational problems. 

3. The warning systems onhoard the Apollo spacecraft and in the Jlission Control Center should be carefully reviered and modified where appropriate. with 

specific attention to the folloiving : 

a. Increasing the differential between master alarm trip levels and expected 

normal operating ranges to avoid unnecessary alarms. 

b. Changing the caution and n-arning system logic to prevent an out-of-limits 

alarm from blocking another alarm when a second quantity in the same snbsystem goes out of limits. c. Establishing a second level of limit sensing in Jlission Control on critical 

quantities \vith a visual or audible alarm wliicli cannot be easily overlooked. 

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

rectctant valves plus a master alarm when any valve closes. 

4. Consuniables and emergency equipment in the LJI and the CJI should 

be reriewed to determine whether steps should be taken to enhance their 

potential for use in a “lifeboat” mode. 

5. The Manned Spacecraft Center should complete the special tests and 

analyses now underway in order to understand 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 JISC 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 rhich 

have previously been corrected or explained. Furthermore, critical decisions 

involving the flightworthiness of subsystems should require the presence and 

Iull 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 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 information developed in this investigation. 

8. SASA should conduct additional research on materials compatibility, ignition, and combustion in strong oxidizers at rarious g levels : and on the characteristics of supercritical fluids. Where appropriate, new SASA design standards should be developed. 

9. The JIanned Spacecraft Center should reassess all Apollo spacecraft subsvstems. and the engineering. organizations responsible for them at JISC and 

a”t its prime contraccors, to huie adequate understanding and control of the 

engineering and manufacturing details of these subsystems at the subcontractor and vendor level. Tl‘here 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. 

COSCLUSIOX 

In concluding, I would stress two points. 

Th first is that in this statement I have attempted to summarize the Board’s 

Report. This Report and its appendices are the result of more than seven weeks of intensive work by the Board, its Panels, and staff, supported by the 

SASb and contractor organizations. In the interest of time, I have not included many supporting findings and determinations which are set forth in 

the Report. 
31 

Tlie second point I wish to make is this: 

The ,ipollo 13 accident, which aborted man’s third mission to explore the 

surface of the moon, is a liar& remindrr of the immense difficulty of this 

undertaking. The total Apollo system of ground complexes, launcli vehicle, 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 success- ful lunar surface explorations is a tribute to those men and women \vho 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, avertecl only by outstanding performance on tlie part of tlie crew and the ground control team which supported them. 

Tlie Board feels that the Apollo 13 acciclent liolds important lessons which. 

when applied to future missions, will contribute to the safety and effectiveness 

of manned space flight. Jfr. Chairman, this concludes my prepared statement. 

Bfr. CORTRIGHT. The prepared statement is a siunmary of the Board 

report, and I will give you a suniniary of the statement. 

As Dr. Paine pointed out, the Reriew Board was created about April 

17, and we first set about to put together a technical and management 

team v-liich could tackle the problem of understanding this particular 

accident. Our charter, in essence, was to find out what happened, why 

it happened, and what do we do about it. 

Sow, the first slide indicates the Review Board that was formed. 

(Slide 1.) 
C D MARRINGION - ' I I PINKEL 

I J E WILSON JR 

APOLLO 13 

REVIEW BOAR

L M CORTRIGHT. CHAIRMAN 

R F ALLNUTT V 1 J

N A ARMSTRONG M KL

J F CLARK H M 

W R HEDRICK - 

1 

1 

J J WILLIAMS 

r -- - 1 

I 1 INCIDENI EVENTS 

I B BALLARD 

Poll INCIDENI 

EVENIS 

H M SCHURMEIER I ____ 

E F BAEHRSUBSYSTEM h 

SYSTEM TESllNG 

RELlABlllIV h 
33 

seated at my left here. Deputy Associate Administrator (Engineering), Office of Space Science and Applications; and Rfr. Milton Klein, 

seated on the other side of Dr. Paine, is Manager of tlie AEC-NASA 

Space Suclear Propulsion Office. 

Dr. Hans Mark, who was unable to be here today, is Director of the 

Ames Research Center. 

As you can see from the organization chart, we put together four 

major panels : Mission events, manufacturing and test, design, and 

project management. Bnd in addition, we have a number of supporting 

staff functions. 

The mission events panel was assigned the job of reviewing in 

meticulous detail everything that telemetry aiid any other type of 

records show as to what happened in the hours preceding the accident, 

actually from liftoff to the accident-that was the preincident events 

portion-then tlie actual few minutes of tlie accident, and then the 

postincident events, which corer the question of how well did we recover from this problem and get the astronauts back to earth. Frank 

Smith from NhSA Headquarters headed up inission events work. 

Manufacturing aiid Test. was headed up by Ah. Schurmeier, whom 

you may remember froin his Ranger and Mariner days. He had three 

groups under him, Fabrication and Acceptance Testing, Subsystem 

and System Testing, and Reliability and Quality Assurance. 

In the area of Design. we brought in Mr. Hinimel, wlio managed 

the Agena aiid Centaur programs. He liad four groups : Design Evaluation, Failure Modes and Xechanisiiis, Electrical, and Related Systems. 

And then we put together a somewhat smaller group, headed by Mr. 

Kilgore, to look into the project nianageinent aspects of this problem. 

Kow, the manner in which we worked at the Manned Spacecraft 

Center is something like this : We relied heavily on the technical teain 

at the Manned Spacecraft Center to generate basic factual data for 

us. In addition, we levied on this group special requirements for 

analyses, additional data, and special tests. 

To do this and to make this a smooth working relationship, lve had 

Ah. Charles Mathews of the Office of Manned Space Flight serve as a 

liaison representative and in addition, we established working relationships which were quite effectiTe and minimized, if not eliminated, 

any duplication of effort. I think this vorked quite effectively. 

In addition, we levied special testing requirements on the Manned 

Spacecraft Center and brought to bear a rather wide NASA effort 

of analyses and special tests. 

For example, at tlie peak of our testing, and I id1 go into that a little bit more later, there were nearly 100 special tests run and several 

hundred people running tests to attempt to duplicate on the ground 

most of the elements of this accident as vie think ve understand it. 

Non-, let me sdch for a moment to tlie problem as it confronted 

us when we first arrived in Houston; of course, the accident had occurred some days prior to that. MTe know from telemetry that the accident probably centered around the No. 2 oxygen tank in the service 

module. I will tell you a little bit more about that tank in LZ minute. 

We did have fairly good although not absolutely complete telemetry. 

That was a strong plus point to start with. TT’e also had the crev 

reports. The crew had heard the bang. At least one member felt a 
34 

shudder. They observed the renting of some forin of material into 

space, and near the eiid of the flight n-lien they separated froin tlie service module they observed that the pinel covering bxy 4 was iiiissiiip. 

Xiicl they brought back photographs of the service moclule froin which 

we attemptecl to glean information about what liacl liappeiied within 

that bag. 

Sow, what tlieii follo\vecl I tliiiik was aptly described by Dr. Paine 

as a rather massive detecting job, and to take you tlirougli that I would 

like to begin by describing the systeiii in which the failure occurrecl 

aiicl place it in the contest of the total -ipollo system. 

The nest slide-of coiirse, -on are all fmiiliar n-itli the basic Apollo 

lauiicli reliicle and where tlie service iiiocliile, the liuiar inodule and the 

command iiiodule are located. The problem was in tlie service module. 

(Slide 2). 

u 

0 

3 m 3 

N 
35 

The next slide (slide 3) gives a little bit inore detail on the sewice 

module. Bay 4 it is callecl sector -1 on this slide. This is the sector in 

which fuel cells, oxygen tanks, ai~cl hydrogen taiiks are located, and I 

will show you that a little bit better in another slide. 

SECTOR 2 

SECTOR 3 

SECTOR 4 

SECTOR 5 

SECTOR 6 

ENVIRONYEUTAL 

CONTROL SUBYSTEY UOZZLE EXTENSION 

RAOIATOR 

FGRKARO B U L K H E A D I N S T A L L 

H Y O R O G E N T A N K S 

SBANO HIGH GAIN ANTENUA 

SERVICE PROPULSION ENGIUE 

SERVICE PROPULSION SUBSYSTEM 1 OXIDIZER TANKS 

OXYGE'I TAkKS HYOROGEN TANKS FUEL CELLS } SERVICE PROPULSION SUBSYSTEM 

FUEL TANKS 

C E N T E R S E C T i O h SERVICE PROPULSION ENGINE AND 

nELlUM TANKS 

Slide 3 

Figure 3-6.- Service module. 

In addition, the propellants for the service moclnles are located 

in adjacent bays and there is a center tunnel here which assumes 

significance in the analysis later on, because as this sector is pressurized by a rupturing tank, the gas flows into this center section and 

pressurizes it. 
36 

If that pressure mere to rise enough, it could have separated the 

command module from the service module-simply blown it off. It 

takes about 10 p.s.i. pressure to do that, whereas it takes something 

in excess of 20 to 25 psi to blow the panel off bay 4. 

The next slide shom in a little bit inore detail the service module 

with the panel removed from bay KO. 4 (slide 4). In the upper shelf 

are the three fuel cells that provide power. They in turii are fed by 

oxygen tanks and hydrogen tanks. 

NASA-S-70-5l2-Y 

ARRANGEMENT OF FUEL CELLS AND CRYOGENIC 

SYSTEMS IN BAY 4 

lide 4 

These are two hydrogen tanks with a cylindrical sleeve to fasten 

them together. The two oxygen tanks sit on this shelf. Oxygen tank 1 is 

here, and KO. 2 is here [indicating]. There is a valve module and servicing panel here. This oxygen tank KO. 2 is the one which failed. 

May we have the next slide, please. (Slide 5.) 
37 

This is the lower section of bay 4 showing the hydrogen tanks and 

the bottom of oxygen tank 2 which failed. All this crinkly material 

which looks a little messy, is designed that way. It is a Iiapton-coated 

Mylar insulation. 
38 

Slide 6 

On the right (slide 6) are shown two of tlie three fuel cells which 

are on the shelf above the oxygen tanks. This is oxygen tank No. 2 

that failed. This is the dome of the tank which covers the spiraling 

tubes which penetrate tlie tank and which carry fluid into and out 

of tlie tank as well as the electrical wiring. 

I might say at this point, JIr. Chairman, that we hare brought 

an oxygen tank of this type with us today, on iny left, which is 

actually one in which n-e rail a full-scale coinbustion test. I will describe that later. 

Kext slide. (Slide 7.) 
39 

Slice i 

This is oxygen tank 2. The lilies run froin the top over to the servicing panel. 
40 

, cap 

A Closeout cap 

Figure 4-1.- Oxygen tank no. ts * 

slide 8 

The next slide (slide 8) is a cutaway drawing to show you vhat is 

inside that tank. Basically, the tank is a double-walled tank. The inner 

wall carries the loads. There is insulation outside of that between 

the inner vall aid outer wall, and this area is pumped to ~OTT vacuum 

to maintain the proper heat leak rates, which have to be very lon-. 

This tank was EI very excellently perforining tank thermodynamically. It is a tough problem to just keep the oxygen that long in space, 

and this was a major development problem in getting this tank built 

in the first place. 

On the right, here, is a heater assembly. There are two heater coils 

wrapped around the tube here, each of which have a thermostatic 

switch, which you will hear more about later. 

The switch 11-as designed to open at 80" F., plus or minus lo", to 

protect the heater assembly froin overheating. At either end is an 

electric motor with a fan. The oxygen is drawn into this tube through 
41 

the sniall hole shown here, it flows downward through the fan and 

is blown outward through these little square holes shown here. 

The fans were required for several purposes actually, one, when 

fans are not used the heat from the heater tends to remain in the 

vicinity of the heater and creates a tliermal bubble. If this gets too big 

and then later on is inixed,.you get a pressure collapse, which is a 

sudden drop in pressure vliich may be undesirable under some conditions. If a thermal bubble exists it interferes with the accuracy of 

the quantity measurements. 

This probe down the middle is a capacitance gage, so-called, which 

measures the density and hence the amount of supercritical oxygen in 

that tank. The oxygen is kept in a supercritical state, which is a single 

phase state. There is never a boundary in it between liquid and gas. 

You can think of it as a very heavy gas. Some people like to think of 

it as a very heavy gas, others as a liquid that never recedes, in a sense, 

it always fills up the rolurne. 

Xow, the fails then break up thi3 bubble when the heaters are on 

rlnd eliminate the stratification, so-called, that takes place there. This 

makes possible accurate quantity measurements throughout the entire 

flight regime. They also make possible the input of larger aniounts of 

heat without the risk of an extra hot region right in here, because 

they stir the contents up and mix it around. 

This tank through the supply line supplies the fuel cells and the 

environnieiital control system. The oxygen goes into pom-er generation 

2nd breathing for the crew. 

There is a pressure slritch on this line, a pressure transducer to 

measure the pressure, and a relief valve in the event this pressure gets 

too high. This relief valve actually opened during the course of the 

accident. 

I believe the next slide (slide 9) shows a little more detail on the 

area diere the problem at least got to its worst stage. 
42 

Figure 4-2.- Oxygen tank wiring and lines. 

Slide 9 

There is a bundle of wires that come up froin the lower fan motor 

and joins with the n-ires from the upper fan motors and the heater 

wires, and they run over through n loop through some holes in this 

Teflon and glass collar. 

There is a temperature sensor mounted right on the collar to measure 

the temperature in the tank. All of this wiring goes up through the 

top of this quantity gage and runs out of the tank through an electrical conduit. It spirals around through this conduit and it is brought 

out here. 
43 

I would like to point out that I am going to refer later to a fill tube 

assembly. This is the fill tube assembly. It consists of three pieces. It 

is possible within the tolerances of manufacture 011 those three parts 

to build that so that it fits in there loosely, and in fact it can fit loosely 

enough so it can be displaced subsequently by either normal handling 

or abnormal handling. I will come back to that. 

I will show you later that we are convinced that a fire started on 

wiring either lower down in the heater assembly or in this region. The 

fire progressed along these wires through the holes in the Teflon element here, probably igniting that Teflon element, setting up a little 

furnace in here which failed at least the conduit, and probably more 

than the electrical conduit. 

A temperature rise of not too much would make it fail at the 1,000 

pounds per square inch pressure it TT’~ then experiencing. Probably 

the penetration cap came out at that time. 

Now, I would like to go on and shorn you the next slide. (Slide 10.) 
44 

This is the best-I won't say the best as projected because you lose 

something in the process of making view graphs, but this is an enhanced photograph taken by the crew of the service module showing 

bay 4. This is the vicinity of the oxygen tank. These are the fuel cells, 

tlie hydrogen tanks over here. There is quite a bit of Mylar projecting 

from this bay which confuses the photography. 

The photograph experts, the photo interpreters who have worked 

days ancl clays on this photography, feel they can find highlights which 

show that the tank is still there. We all believe that the tank is there. 

At least half of the people-probably more-who look at it can? see 

it. It turns out-and x-e have duplicated this in the laboratory-that 

with tlie particular geometry of the tailks aiicl with that lighting condition, it is l-ery liarcl to see in any event. Certainly, we were not able 

to deterniiiie the coiiditioii of the tank, so the main value of this 

photograph, I feel, is tlie coiiclusire ericlence that the panel is gone. 

If )ye hadn't had this observation, I am not sure that we would have 

. 

concluded that. 

I would like to go to the next slide (slide 11) to go back to the 

telemetry. 

. 
45 

47-S91 0-70-4 
46 

Could I have the next one (slide 12) ? 

& - 

LL 

LA 

c 

CL 

U 

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lA 

c3 

W 

vr W 

z 

m 

W 

E 

x 

Id 

W 

a 

W IE 

vr + 

W 

I 

I3 

0 

LA 

a 

E 

a 

t 

E 

W Iz 

Y 

E 

LAz 

d 

W 

CL 

W 

LA 

2 

W 

z 

z 

m 

W 2 

0 

m 

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LA 

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CL 

2 

W 

CL 

The slide on the right (slide 11) is a portion of the telemetry record, 

and I have extracted events from that record and listed them in an 

easier manner to follow the slide on the left. 

Here are some voltage spikes on a system, an attitude control system 

that was connected to the same electrical bus as the tank on which me 

had the problem. This was an indication of something happening in 

the tank. These are referred to as glitches. 

These are two large current spikes which occurred here. They are 

both short circuits. The first one is the one which we now feel started 

the fire in the tank. It occurred immediately after turn on of the 

fans. 
47 

Shortly thereafter, the pressure in the tank began to rise and you 

can see this pressure rise here up to about a value of 1,008 pounds per 

square inch. Later on, I will refer to tests that have duplicated in 

general this condition. 

The temperature indication rose much later but, if you recall from 

the last slide I had on, the temperature sensor is located on the quantity probe. 

Here is the quantity probe. Here is the temperature measuring 

device. You wouldn't expect that to measure a temperature increase 

until the fire progressed along the wire to that vicinity. That is why 

the temperature rise is delayed. Everything let go about this point. 

There was a dropout of telemetry and later on a decrease of the oxygen pressure in tank No. 1, which was also failed by this rupture. 

There are the events summarized here. Fans on, short circuit starts 

the fire, the pressure begins to rise, and this pressure rose, by the way, 

because as the oxygen is heated by the fire it expands in sort of a bubble 

around the wire and compresses the rest of the oxygen in the tank. 

Temperature begins to rise IT-hen the fire gets to the vicinity of the 

temperature gage. The highest pressure reading occurred here. 

We later ran tests which showed that the pressure dropped off at 

that point because the relief valve opened and it was determined that 

the relief valve was sufficient to drop the pressure at the observed 

rate. 

The panel separated here. The telemetry dropped out. The crew 

reported the problem. Aid very soon thereafter fuel cell 3 and fuel 

cell 1 failed, because the supply valves had closed from the shock of 

the explosion and the blowoff of the panel and thus shut off the fuel 

cells. 

The CHAIRJIAN. Vhat vas about the elapsed time from the time 

you got the first spike? What IT-as the time elapse between the first and 

second spikes 1 

J4r. CORTRIGHT.hese were different short circuits. The fire had already begun and in consuming insulation could have contributed 

to the second one. 

The CHAIRMAS. Aren't the lovier figures the time element ? 

Mr. CORTRIGHT. Yes, hours and minutes. 

The Cmmxn-. That I can see, I as looking at the other chart. 

&. CoRmrGHT. That is hours, minutes, and seconds. The scales are 

different. This goes from 55 hours 53 minutes and 15 seconds to 55 

hours, 55 minutes, over here, if my eyes aren't failing me-so there 

is 5 seconds between each major block. 

The CHAIRMAN. I see. 

AIr. FTTLTOS. Is your graph in real time ? 

JIr. CORTRIGHT. Yes. 

Mr. FIXTOS. That is the real time graph? 

llr. CORTRIGHT. Yes, it is an accurate representation of what happened. 

Kow. having this telemetry and the creiT reports, it didn't take us 

too long to reach the conclusions I have given you so far. We then 

decided it was time to prove that this analysis vas correct, so we had 

to ask ourselves a series of questions and then proceed to ansver them 

and prove our answers. 
48 

The first question was : Was combustion required to raise the pressure the indicated amount ? Aiialyses showed that you could not get 

increases of that amount simply by feeding electrical energy alone 

into the tank as a heat source. There was no other possible heat source, 

so the coiiclusion was that yes. combustion was required. 

The next question then is: 7Vhat was there in the tank that could 

burn? The materials in the tank had passed characteristics of materials 

of the "COJLQT" system which is used to determine acceptability of 

materials in oxygen environments. 

We made an investigation and determined that Teflon can burn 

under these supercritical oxygen conditions, as can aluminum. solder. 

and other materials present. 

Jlr. HECHLER. Atx-hat temperature does Teflon burn? 

Jlr. CORTRIGHT. The actual temperature of combustion are over 

2.000' but ipiiition energy required is as lon- as one joule if it is done 

through an electric arc. 

Jlr. PRICE. Why did not the relief valve take care of that? 

Jfr. CORTRIGHT. When the pressure built up due to the combustion, 

the relief 1-alre held the pressure from exceeding about 1.008 pounds 

per squnre inch. At the same time, the fire progressed up to the metals 

of the tank wall and its tubing and overheated them. at which poifi 

they lost their strength and failed. That was the manner of the failure. 

Mr. PRICE. It wasn't pressure in the tank that blew the tank? 

Mr. CORTRIGHT. In a way it was. The tank had plenty of strength. 

It was twice as strong as necessary to hold a thousand pounds per 

square inch at cryogenic temperatures, but when it was heated by 

the fire it lost its strength and failed locally. So it is a combination 

of effects. The material loses its strength and the pressure blows 

through the weakened portion. 

Mr. DOWXING. At \That time on the time scale did the explosion 

occur 1! 

Mr. CORTRIGHT. Right here where you see pressure drop. 

Mr. DOWNING. Transpose that over to the other chart. 

Mr. CORTRIGHT. Right here, panel separates, telemetry drops out. All 

of this occurred in a very brief period of time. 

I have not expanded the scale enough to break the details down into 

milliseconds. 

All right, so getting back to the questions, we had to ask ourselves 

what would burn, and I just told yon what would burn. 

Then we had to ask ourselves, well, how could you start these materials burning since generally speaking they are considered compatible with oxygen, and I think here me ran into a new phenomenon that 

was not recognized or widely recognized before. and that is that Teflon 

can be ignited rather easily if an electric arc is the igniting mechanism, 

and the combustion will propagate in supercritical oxygen. We ran 

tests to show that these small amounts of electrical energy were sufficient if they were in the form of an electrical arc which concentrates 

the heat very locally in the material, and we ran additional tests to 

show that even through the 1 amp, relatively quick blow fuse that 

was on that line to protect it, you could get energy 10 to 100 times in 

excess of what was required to ignite the wire insulation. 

Now, having determined that, we said to ourselves, well, if we had 

that type of fire mould that be consistent with the times of the pressure 
49 

and temperature rise? So we made measurements of the rate of combustion of Teflon wire insulation in supercritical oxygen, and I will 

shorn you photographs of this later but, briefly, we found that the rate 

of burning varies at 1 g., because, just like the smoke from the candle 

goes up if you are burning it, more fresh oxygen is drawn in by convection and feeds it. 

If the fire is burning up, it goes fast; horizontally, medium, and 

down, slow. When it is burning down, it burns about one-quarter of an 

inch a second. 

We ran tests at zero gravity and found it burns from one-eight to 

one-quarter of an inch a second at zero g., so it burns still more slowly. 

Going back to the lengths that were in the tank, we were able to confirm that there was correlation here between what we thought would 

happen and what happened, and later on then we took an actual tank, 

the one sitting over here, and ignited its wiring insulation and recorded its temperature and pressure history. 

Mr. FULTON. Mr. Cortright, the question of burning insulation 

through arcing has come up earlier in our previous Apollo investigation, and also the question of what kind of insulation. 

Some of us went down to Houston and various places and were 

surprised, amazed and shocked at the way the insulation on wire 

burned after there had been an arcing situation. It looked to us as 

if it were a sparkler on the Fourth of July. You could see it just 

running along the mire emitting these sparks. 

Kow, when we had the Apollo 204 insulation question, had this 

material been tested for electrical arcing, or was this a new situation 

that even the manufacturer had not considered in spite of Apollo 2041 

Mr. CORTRIGHT. The testing that followed Apollo 204 concentrated 

on the environment of the crew quarters, which is a much lower pressure. The testing was not extended to include the conditions within 

the oxygen tanks themselves, where supercritical oxygen is stored. 

Mr. FCLTON. That is my point. When we had much lower pressure 

in a crew cabin atmosphere and ve could see an qrc and the insulation 

burning with sparks being emitted, why then, with this oxygen situation and a much higher pressure and possible arcing, wasn’t that 

gone into? 

Mr. CORTRIGHT. Mr. Fulton, the wire that caused the problem in the 

204 was a polyvinyl chloride, I believe, and we switched to Teflon 

and Teflon was qualified in the pure oxygen and the polyvinyl chloride 

was totally eliminated. 

Mr. Fnms. It is no longer qualified? 

Rfr. CORTRIGHT. Teflon is qualified for the cabin, to the best of our 

knowledge. It is definitely combustible at supercritical conditions in 

high pressure oxygen. 

Mr. Fnmx. Do you need, in your estimation, further tests along 

these lines of combustibility as well as insulation subjected to arcing 

conditions? Is this just one stage, or should me go much further 

before we say that the hookup is safe? 

Rlr. CORTRIGHT. We have a recommendation along the lines you are 

suggesting, to conduct considerable additional research on this problem within NASA and to revise our standards where appropriate. 

I think more work has to be done, because we have learned so much 
50 

in the past 2 months that we tell ourselves we should have known 

earlier. 

JIr. FULTON. Tas this covered by any technical report of NASA, 

this particular arcing situation in respect to insulation under high 

pressure in an oxygen atiiiospliere ? 

JIr. CoRrKIGI-IT. Sot to my kno\Tledge. 

JIr. FL-LTOS. Why wasn‘t it ? 

Mr. CORTRIGHT.eflon was felt to be safe under these conditions. 

In hindsight. if we had conducted research on various manners in 

which Teflon might burn in this very high pressure environment, we 

~~onld have fonncl out it can be, but this was not clone. 

JIr. FTTLTON. Did XA4Sd accept the statement of industry or the 

industry catalog or manual on the characteristics of Teflon under 

arcing conditions, or did it do its ovn separate investigation on it? 

JIr. CORTRIGHT.o answer your question as best I can, I am quite 

certain I nin not familiar with all the governing documentation or 

arailable literature on the subject. The Teflon was qualified by means 

of an iinimct test which is one method n-idely used for determining 

compatibility of materials in all types of oxygen environments. 

This test consists of a blo~ on the material and sensitive materials 

c~n be ignited in this manner. If they dl survive a blov of a certain 

intensity. they are considered compatible. 

Mr. FI-LTOX. At That point does that responsibility rest ? With 

SATASAl ? The contractor ? The manufacturer ? Where does that responsibility rest ? Certainly not \Tit11 the contractor. 

hlr. CORTRIGHT. I think the responsibility for seeing that our equipment is flightworthy rests with SASh. 

The CIIAIRJIAX. This might haTe been unanticipated because in the 

first place there haven‘t been a lot of places in SASA or in industry 

where you would duplicate such conditions under high oxygen pressure. 

hlr. CORTRIGHT. This is the only case we know of, Mr. Chairman. 

There map be others. It is the only case we know of \There wire insulation of this type as ignited and burned in an environment like this. 

Teflon is generally considered to be the best of the flexible insulations available. 

The CI-IAIRMAX. hiid of course we have had precedence in this at 

NASA n-here re found that the generally accepted types of welding, 

for instance, in the case of the Centaur. proved that they vere not fit 

and they had to go out and do it over again. This is part of the progress 

that we make in the space effort, is it not? 

Mr. CORTRIGHT. That is true. 

Mr. HECHLER. Does your recommended research entail use of different material other than Teflon? Are there other materials? 

Mr. CORTRIGHT. Yes, sir. The recomineiidatioii corers the compatibility of various materials with pure oxygen under other conditions, 

further research into supercritical oxygen and examination into other 

modes of propagation. 

hfr. HECHLER. As a layman, I didn’t understand your use of the 

term “one joule.” 

Mr. C~RTRIGHT. That is a measure of electrical energy which would 

go into a spark. If you heat Teflon, it depends on how you heat it. If 

* 
51 

you just heat it in an oxygen environment, the first thing that will 

happen is that the insulation will start to deteriorate. 

I am going to give you a roundabout answer to your question and we 

have run these tests subsequent to tlie accident and actually, as I will 

shorn later, that did happen to this wiring before tlie launch. But 

when it gets up around 800°F. or 900°F., the insulation will sloKly 

oxidize away and disappear of7 the wires entirely. 

If you, say, ignite a local portion by means of a Sichrome wire, 

which would be a glowing white hot n-ire, m-hen the Teflon gets to 

1,300"F. in high density oxygen, it will react in a combustible manner, 1,300"F. is the figure you were asking for. It then burns at a higher temperature. 

Mr. FULTON. Why was this in the oxygen tank when it looks as if 

there was room outside? 

Mr. CORTRIGHT. We need heaters in tlie tank to keep the pressure 

up, to keep the oxygen feeding properly into the fuel cells and into 

the crew compartment. The fans were required to stir the contents. 

Now, the particular mechanization that was used was two electric 

motors in the tank, and n-e cover that in our report. 

Jlr. FCLTON. Why was everything put inside the tank when there 

was some danger of arcing? 

Mr. CORTRIGHT. I think that is a fair question. This particular de- . sign approach chose to do it that way. What was done at tlie manufacturer's plant is that very meticulous assembly procedures n-ere developed to prevent or minimize the chance of damage to the electrical 

wiring so that short circuiting could not or probably would not take 

place later. 

We think that the design was deficient. It turns out that the basic 

design wasn't really at fault in this case, the wire as damaged by 

an overheating condition that I am going to describe for you. 

Jlr. FGLTOS. In sumniing up the basis of the opinion which you 

developed, the gentleman from West Virginia and the chairman and 

I think you were citing the adrance state of art not only on the material but on the circumstances that had developed in the manufacture; 

is that your opinion ? 

Mr. CORTRIGHT. Jlr. Fulton, I believe that the tank when itwas built 

constituted a very advanced tank, and we know much more today about 

the sorts of problems you can pet into with a tank of this type, and also 

other ways in which it might be designed and put together that viould 

make a more reliable unit out of it. 

I think that is what you told me, and I agree with it. 

Mr. FULTOX. I say it is not any negligence or any failure either on 

the person selecting the material, the insulation, the engineering, the 

design, or the operation of the ~~ehicle-it is rather that a combination 

of circumstances created a requirement for a new advance in the art 

which you are now doing. Is that it? 

Mr. CORTRIGHT. Not entirely. I believe there were deficiencies in the 

tank design and the manlier in iThich the tank was handled, and I will 

go into that in tlie balance of my statement. 

Mr. HECHLER. I also would like to observe, and I will develop it 

further, I think we need an advance in the art of administration as 

well as technical design, but I will develop that later. 
52 

Mr. PRICE. I would like to ask you, do you think that NA4SA's qualification of material that is assembled into our Apollo equipment is 

sufficient, or do you think it could be improved, not just in this case 

but in the overall situation? 

Mr. CORTRIGHT. Yes, I think we can improve and we are always 

trying to. 

In this particular case, or in all cases actually, there are many safeguards to insure that nonflightworthy hardware is weaned out. Sometimes due to an unusual combination of events or mistakes, this does 

not happen. The Agency has a meticulous system to prevent this from 

happening. No system is perfect. Problems can sneak through. 

Mr. PRICE. What progress has been made in this particular area? 

Has this been altered so \ye can continue with the next flight, or is 

considerable change necessary ? 

Mr. CORTRIGHT. The hardware itself is being redesigned for modification at this time, aiid within the next few weeks I expect that selection of the design will be made. 

As far as the procedures are concerned, I think that every element of 

the organization is taking another look at the procedures it has been 

using in the light of what we have learned here, to make sure it doesn't 

happen again. 

Mr. PRICE. When you accept parts like this, do they, for instanceflush the system or try the system before it is put in? Is each individual 

part as it comes from the company checked? Do we have controls 

in accepting the parts ? 

Mr. CORTRIGHT. Depending upon the particular compound or parts, 

there are flight acceptance test specifications which normally a subsystem would have to meet. The individual parts may or may not 

have acceptance testing depending upon what the part is, but they 

have to meet certain standards. 

The specifications for these parts are written at various levels, 

some are written by the subcontractor to the vendor, others are written 

by the prime contractor to the subcontractor, and some top level 

requirements are written by KLiSA to the prime contractor, so there 

are various levels of checks aiid balances and rei*ieF at work here. 

Mr. PRICE. no you think we should go further and go into these 

companies with SA4Sd's own inspectors and where feasible run a 

test that will meet that standard-in other words, the company will 

say yes, we have met your requirements, then is it tried out to see 

if it meets requirements? Is this done? 

Mr. CORTRIGHT. This vas normally done. In this particular case 

there was a thermal switch which was not tested, and this came back 

to bite us. Normally what you asked us for is done. 7T'e are all reviewing our procedures. 

Jlr. PRICE. Whose responsibility was it ? 

Mr. CORTRIGHT.\TA4SA accepts responsibility for the total system. 

This cannot be delepted. 

Mr. PRICE. When it was built, whose responsibility was it ? 

Mr. CORTRIGHT. The prime contractor and subcontractor who manufactured this equipment had a certain responsibility to insure that 

the testing was sufficient. I will come to that in more detail later. 

Mr. A~OSHER. Mr. Cortright, this thermostatic switch, was that same 

equipment on A4pollo 11 and on Apollo 122 
53 

Mr. CORTRIGHT. Yes ; aiid on the earlier Spollo flights. 

Mr. RIO~IIER. What then was the crucial difference or the crucial 

event that inacle the difference between those flights and this flight 1 

Jlr. CORTRIGHT. The crucial difference TTXS a special detanking procedure that took place at the Cape, and I will go into that later. 

I am delighted to answer all the questions I can, but it might make 

a. more coherent story if I quickly finish up. It will only take me 10 

minutes. 

Mr. KARTH (presiding). With tlie indulgence of the committee, I 

think we should wait to ask questions until he finishes. 

JIr. CORTRIGHT. Thank you. 

There is a coherence to this that hasn't become apparent yet; I 

lioiie it ~dl. [Laughter.] 

I was going through the problems that we faced aiid the things we 

have to prove to ourselves. I believe I had gotten down to the point of 

why the-I had explained that the pressure rise history was consistent 

with the rates of burning along the wire, and we proved that by test. 

We also postulated that the temperature rise delay could reasoltably be expected to occur with fire remote from tlie sensor, and this 

was later demonstrated in tests. 

The maximum pressure correlated il-ith tlie relief valve operation. 

The telemetry dropout was the next question. Why did we lose 

telemetry ? 

TKO things happened. One was that there WLS a strong shock to the 

spacecraft at the time the panel blew off, and second, from the photograph, the high-gain antenna on the service module was bent. 

It seems reasonable that either parts of the panel coming off, which 

you will see in a motion picture of a model test, would have done that, 

or some other part coining out of that bay. 

The loss of pressure from oxygen tank No. 1 is surmised to have 

occurred from one of two causes, either shocking open of the valve 

or cracking of one of the high pressure lines from the tank. 

Now, having put those pieces together and run the tests to validate 

them, the question was.does it all hang together and make a coherent 

story ? 

As I may have mentioned earlier, aiid I will repeat it for emphasis, 

we did bring the total resources of the Agency to bear on the investigation which involved all of our centers, plus the prime contractor, 

North Ainerican Rockv-ell, the subcontractor who delivered these 

tanks to Sortli American Rockn-ell, Beech Aircraft Corp., and a number of other companies helped. 

This included about 100 special tests involving several hundred 

people and I would like to sholy you a film now which gives you highlights of this testing program. 

(Film shown.) 

Rfr. CORTRIGHT. The first thing we had to do as to demonstrate 

that we could ignite Teflon with-the lower energies that were available. This was a test done in Houston at the Manned Spacecraft Center. An electric arc at the left ignites the Teflon wire, which burns 

along the wire toward the right. It burned out another wire, and 

now it will progress across. 

This is burning in the very high pressiwe supercritical oxygen oi 

the type used in tank KO. 2. These photographs are taken at normal 

speed. It burns along like a fuse. 
54 

This is combustion taking place in iz test rig, at the Lewis Research 

Center, which later on will be used to demonstrate zero p. combustion. 

This is one g., to give you a point of comparison. The smoke is going 

up. At zero g., smoke doesn't know enough to go up because there is 

no up. 

The temperatures of supercritical oxygen for tests like this range 

from minus 100 to minus 200" F. This is burning in an extremely 

cold environment. 

Here is the same test run at zero g., first to illustrate the type of rig. 

this container is dropped 500 feet into a silo where it impacts into 

plastic spheres to absorb the shock. This is combustion at zero g. at 

a rate about half that that occurs at one g. 

These photographs are taken at 400 frames a second because the 

entire time of the test was 8 to 4 seconds. The apparent out-of-focus 

nature of it is caused by the refraction of the supercritical oxygen, 

not by tlie camera. 

We also ran tests to sliow that wire that had been baked at high temperature burned similarly. This is a bunclle that is being ignited in a 

simulated tank. It will burn through the wires down into the Teflon 

collar. It ignited tlie Teflon collar and burned a 2-inch hole at the top. 

This is more of a boilerplate tank. The rupture comes riglit through 

here, very rapidly. There is the rupture. 

hlr. FULTOS. Where does the arcing occur on those pictures? 

Air. CORTRIGHT. 1 Kill come back to that, Ah. Fdton, if that dl 

satisfy you. 

Now, this is moving throueh n fnll-scale tank. the one that I have 

with me this morning: The fi; is burning inside-it just blew through. 

This is an escaping mixture of gaseous and liquid oxygen which took 

place in that rather confined compartment of bay So. 4 of the 

spacecraft. 

Mr. Fulton, to come back to your question, the last test you saw, 

ignition was achieved either with a Kichrome wire or a squib rather 

tlian an electric arc. The electric arc tests were a separate series of tests 

that rTe ran. 

Depending upon the size of the hole, it may be necessary to pet additional pressure in the bay to get it off. These tests were run at Lanpley 

to demonstrate that the oxygen products accompanied by sparks and 

burning material from the tank are sufficieiit to ignite tlie Mylar iiisulation tliat you saw earlier fill the bay. 

TVe have measured augmented pressure rises as much as a factor 

of a hole in the tank between 1 and 2 inches in diameter supplemented 

by this combustion which has been demonstrated, would be sufficient 

to take the panel off the tank. 

This particular test is a sloir- motion film of oxygen combustion 

with tlie Mylar. At the same time, we ran analyses to determine what 

type of pressure pulse would be required to take a panel off, and then 

ran tests which you are about to see to measure, in fact, whether our 

calculated pulses would in fact take the pxnel off. 

Here is a film taken at 2.000 frames a second, and this panel is n 

half scale honeycomb panel blowing off witli n simulated tank rupture, the tank being in this location. That all takes place in n few 

milliseconds. 

It can occur in such a way that tlie pressure builds up highest right 

in this vicinity and to a lesser extent in the rest of the bay and to a still 

of 6, it probably wouldn't be quite that in high flight. A combin a t' 1011 
55 

lesser exteiit in t,he rest of tlie service iiiodel, so that it would not have 

blown off the command model. 

Tlie combiiiatioii of analyses and tests in this point of time provide a 

f airly good reproduction of TThat probably hnppeaed. 

That concludes the film. 

At this point, Mr. Chairman, I would like to wrnp up That I have 

been telling you for tlie past 45 minutes or so, by reading a portion 

of the A4pollo 13 Review Board report. I am turning to the report here 

because these words are carefully chosen aiid I think words should be 

carefully chosen when we come right down to the point of what most 

likely happened aiicl what roles various orgniiizatioiis played in t,he 

problem. 

In reachimng 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 tlie result 

of a chance malfunction in a statistical sense, but rather resulted from an unusual combination of mistakes, coupled with a someTyhat deficient and un- forgiving design. In brief. this is what happened : 

a. After assembly and acceptance testing, the oxygen tank NO. 2 which flew on Apollo 13 was shipped from Beech Aircraft Corporation to I\’orth American 

Rockwell (NR) in apparently satisfactory condition. 

21. It is now known, however, that the tank contained two protective thermostatic switches on the heater assembly, which were inadequate and mould 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. 

d. In itself, the displaced fill tube assembly was not particularly serious, but 

it led to the use of improvised (letanking procedures at RSC which almost cer- tainly set the stage for the accident. c. .Uthongh Beech did not encounter any problem in detanking during acceptance tests, it was not possible to detank oxygen tank No. 2 using normal procedures at KSC. Tests and analyses indicate that this was due to gas leakage 

through the displaced fill tube assembly. 

f. The special detanking procedures at KSC subjected the tank to an extended 

period of-actually about 8 hours-heater operation and pressure cycling for 

about 2 hours. These procedures had not been used before, and the tank had not 

been qualified by test for the conditions experienced. Hoivever. the procedures 

did not violate the specifications which governed the operation of the heaters at 

ESC. 

8. In revieving these )procedures ‘before the flight, officials of SASh, SR. and 

Beech did not recognize the possibility of damage clue to overheating. JIany of 

theae officials were not nware of the extended heater operation. In any event. 

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

lt. A nunlber of factors contributed to the presence of inadequate therinostatic 

switches in the heater assembly. The original 1962 specifications from NR to 

Beech Aircraft Corporation for the tank and heater assembly specified the use of 

28 V clc Don-er, which is used in the spacecraft. In 1%. StR issued a revised 

specification which stated that the heaters slionld use a 63 V dc power supply for tank pressurization : this n-as 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. 

Jfr. FTTLTON. ~~onld you sny that, again? 

Jir. CORTFXGHT (continuing) : 

Beech ordered switches for the Block I1 tanlrs but did not change the switch 

specifications to be coinpatible with 65 V dc. 

i. The thermostatic switch discrepancy was not detected by NASA, NR, or Beech in their revier of documentation, nor did tests identify the incompatibility 

of the switches with the ground support equipment (GSE) at KSC, since neither 

qiialificatioii nor acceptance testing required switch cycling under load as should 

have been done. It was a serious oversight in which all parties shared. 
56 

j. The thermostatic switches could accommodate the 65 V dc during tank pres- surization because they normally remained cool and closed. However, they could 

not open without damage with 63 V 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 inoperative as protectire thermostats. 

By the way, I do have a failed switch here, one from the test, which 

I can show you later. 

12. Failure of the thermostatic switches to open could have been detected at 

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

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

temperature limit and switch opening should hare 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 8-hour period of heater operation. Such 

heating has been shown hy 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 90. 2 n-as in a hazardous condition 

when filled with oxygen and electrically powered. 

Just to digress for a moment, I would like to show my last three 

view-graphs (slides 13, 14, and 15) which shoF the manner in which 

the contacts of the thermostatic switch can weld together. 

Slide 13 
57 
58 

Here are the welded together electrical contacts of the thermostatic 

switch when subjected to about 11/2 amperes, all that it will carry at 

the 65 volt d.c. This is the condition of the vire when removed from 

a test heater assembly. These wires run up through a conduit inside 

the heater assembly, which runs from the other side of the tube from 

the electrical heater. This was liquid nitrogen. ,111 you see is thermal 

damage. In oxygen, one would have expected some of that to slowly 

oxidize away. In some cases it can totally disappear. 

m. It was not until nearly 56 hours into the mission, however, that the fan 

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

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

oxygen tank probably overheated and failed the wiring conduit where it en- ters the tank, and possibly a portion of the tank itself. 
59 

a, The rapid expulsion of high-pressure oxygen which followed, Possibly auglnented by combustion of insulation in the space surrounding the tank, 

off the outer panel to bay 4 of the SlI, caused a leak in the high-pressure SYStern 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 

wrformance on the part of the crew. Nission Control, and other members of 

the team which supported the operations successfully returned the crew to 

Earth. 

large amount of material is included in our report and in Appendix B (Report of the Apollo 13 Review Board) to show the 

manlier in vliich the mission control and the crew coped with this 

in-flight emergency. I think it was truly admirablle. I would commend 

it to your reading. I have not taken time to go through that this 

morning. 

I would say this also: In investigating the accident to Apollo 13, 

the Board has also attempted to identify those additional technical 

aiid management lessons which can be applied to help assure the success of future space flight missions ; several recommendations of this 

nature are included. 

I will no^, on behalf of the Board, state that we recognize our report 

5s being preoccu ied with deficiencies, that is the nature of a review 

board. We feel t f iat the deficiencies we hare uncovered will help the 

program to do a better job in the future, and that they should be 

vieJI-ecl in the light of the considerable successes that this equipment 

and the people who build aiid operate it have achieved today. 

JIr. Chairman, I now vould like to read the recommendations of 

the Board, and this will conclude my statement. This is on page 540 of 

the summary report of the hpollo 13 Reriew Board, this particular 

volume, if you n-ant to read along with me : 

1. The cryogenic oxygen storage system in the service module should be 

modified to : 

(a) Remove from contact with the oxygen all wiring, and the unsealed mo- tors, which can potentially short circuit and ignite adjacent materialIncidentally, page 540 is the rery last thing in the volume. 

Jlr. KARTH. Page 42 in the copies of your statement the members 

have ? 

Ah. CORTRIGHT. Itis the statement on 42, it is in the report at page 

540. They both say the same thing. 

(a) Remore from contact with the oxygen a11 wiring, and the unsealed motors, 

which can potentially short circuit and ignite adjacent materials ; or otherwise 

insure against a catastrophic electrically induced fire in the tank. 

(b) Minimize the use of Teflon. aluminum, 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 opera- tional problems. 

3. The n-arning 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 : 

((I) Increasillg the differential between master alarm trip levels and expected 

normal operating ranges to aToid unnecessary alarms. 

(b) Changing the caution and warning 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. 

(c) Establishilig a second level of limit sensing in Mission Control on critical 

quantities with a visual or audible alarm which cannot be easily overlooked. 
60 

((2) Providing independent talkback indicators for each of the six fuel cell 

reactant valves plus a master alarm n-hen any valve closes. 

4. Consumables and emergency equipment in the LM and the CM should be 

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

for use in a "lifehoat" mode. 

5. The JIanned Spacecraft Center should coinplete the special tests and analyses now underway in order to understand more completely the details of the hpollo 

13 accident. In addition, the lunar module power system anomalies should receive 

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

the areas of analysis and test. 

6. N'henerer significant anomalies occur in critical subsystems during final 

weparation for launch. standard procedures should require a presentation of all 

prior anoinalies on that particular piece of equipment, including those which have 

previously been corrected or explained. Furthermore, critical decisions inrolving 

the flghtworthiiiesr of subsystems rhould require the presence and full participation of an espert who is intimately familiar with the details of that 

subsystem. 

7. SASA should conduct a thorough reexamination of all its spacecraft, launch 

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

osidizers, to identify and evaluate potential combustion hazards in the light of inforination developed in this investigation. 

8. NASA should conduct additional research on materials compatibility. ignition, and combustion in strong oxidizers at rarious g levels ; and on the characteristics of supercritical fluids. Where appropriate, new SASA design standards 

should be dereloped. 

9. The Manned Spacecraft Center should reassess all Apollo spacecraft subsystems, and the engineering organizations responsible for then1 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. 

J4r. Chairinan. that coinpletes my present a t' 1011. 

JIr. R.wnr. Thank you very much, Mr. Cortripht, for your suininary 

report. Congratulations to you and the Review Board are in order for 

haring made what I consider to be a rery positive, definitive and candid analysis of the accident. 

Certainly it gires me confidence that an in-house investiption can 

be made, which in the final analysis can result in criticism if the situation merits criticism, so I want to congratulate you and the Board. 

I am going to ask the members to adhere to the 5-minute rule to 

give everyone on the committee an opportunity to ask questions. 

The Chair recognizes Rfr. Fulton. 

hfr. FULTON. I believe you have done a careful and excellent job on 

the Review Board. 

I would like, with the chairman's permission, to have the recommendations of the Review Board of Apollo 204 put immediately after the 

recommendations that have just been made by this Apollo 13 Review 

Board. 

Air. I~ARTH. Without objection, it is so ordered. 

(The recommendations of the Review Board of Apollo 204 are as 

follom :) 

BOARD FIRDINGS, DETERMINATIONS AND RECOMMEKDATIONS 

In this Review, the Board adhered to the TJrinciple that reliability of the 

Command Module and the entire syeteni involved iii its operation is a requirement common to both safety and mission succec4. Once the Comniand Module has 

left the earth's environinent the occupants are totally dependent upion it for their 

safety. It follows that protection from fire as a hazard involves much more 
61 

than quick egress. The latter has merit only during test periods on earth when 

the Colmiand Jloclnle is being readied for its nlimssinn and no't during the mission 

itself. The risk of fire must be faced; however, 'that risk is only one factor 

yertainillg to the reliability of the Command Module that must receive adequa te consideration. Design featnres and operating procedures that are in'tencled 

to reduce the fire risk must not introduce other serious risks to mission success and safety. 1. FIh-DING 

(a) There was a momentary poTrer failure at 23 :30 :W GMT. 

( b j A detailed design review be conducted on ,the entire spacecraft communica- 

(c) So biiigle ignition source of the fire was conclusively identified. 

Dcterni in,ution 

The most probable initiator was an electrical arc in the sector between the 

--T and +Z spacecraft axes. The exact location best fitting the total arailable 

information is near the floor in the lsoiver forward section of the left-hand 

equiyment bay where Enr~ironnlental Control System (ECS j instrumentation 

power wiring leads into the area between the Environmental Control Unit (EC'U) 

and the oxygen panel. So evidence was discovered that suggested sabotage. 

2. FISDISG 

(a) The Conimand Module contained many types and classes of combustible 

(a) The test mas conducted with a 16.7 pounds per square inch absolute, 100 

material in areas contiguous to possible ignition sources. 

percent oxygen atmosphere. 

D cfo'm iii a t ion 

Rcconi iii ofrlatio?i 

must be severely restricted and controlled. 

The test conditions were extremely hazardous. 

The amount and location of combustible materials in the Command Xodule 

3. FISDISG 

(CL) The rapid spreacl of fire caused an increase in pressure and temperature n-hich resulted in rupture of the Conimand Jlodnle and creation of a toxic 

atmoslihere. Denth of the crew n-tis from asghrxia due to inhalation of toxic 

gases due to fire. A contributorF cause of death was thermal burns. 

( b f Son-nnifonn distribution of carboxyhemoglobin was found by autopsy. 

Dctcrnzincltion 

Autopsy data leads to the medical opinion that unconsciousness occurred 

rapidly and that death followed soon thereafter. 

4. FISDISG 

Due to internal pressure. the Command Module inner hatch could not be 

oiwned prior to rupture of the Command JIodnle. 

Dctcrizi inaf ion 

The crew was never caliable of effecting emergency egress Wanse of the 

pre.wirization before rupture and their loss of consciousness so011 after rupture. 

Rccom ni cnda tion. 

The time reqnired for egress of the cren- be reduced and the operations neces- saqr for egre.Gs be simplified. 

5 FIXDING 

Those organizations responsible for the planning, conduct and safety of this 

test failed to identify it as being hazardous. Contingency preparations to permit e-cape or rescue of the crew from ail internal Command Module fire were not 

made. 

(a) No procediires for this type of emergency had been established either for 

the creiv or for the spacecraft pad work team. 

(b) The emergency equilinient located in the White Room and on the spacecraft n-orli levels was not designed for the smoke condition resulting from a fire of this nature. 

47-591 0-70-5 
62 

(c) Emergency fire, reecue and medical teams were not in attendance. 

(d) Both the spacecraft nwrk levels and the umbilical cower accePs arm con- tain features such as steps, sliding doors and sharp turns in the egress paths vhich hinder emergency operations. 

Determination 

Adequate safety precautions were neither established nor observed for this test. 

Recommendations 

(a) Management continually monitor the safety of all test operations ana assure the adequacy of emergency procedures. 

(b ) All emergency equipment (breathing apparatus, protective clothing, deluge systems, access arm, etc. ) be reviewed for adequacy. 

(c) Personnel training and practice for emergency procedures be given on a regular basis and reviewed prior to the conduct of a hazardous operation. 

(d) Service structures and umbilical towers be modified to facilitate emergency operations. 6. FIXDING 

Frequent interruptions and failures had been experienced in the overall com- munication system during the operations preceding the accident. 

Determination 

The overall communication system was unsatisfactory. 

Recommendations 

(a) The Ground Communication System be improved to assure reliable com- munications between all tests elements as soon as possible and before the next 

manned flight. 

(b) A detailed design review be conducted on the entire spacecraft communication system. 7. FINDING 

(a) Revisions to the Operational Checkout Procedure for the test were issued at 

5 :30 pm EST January 26, 1967 (209 pages) and 10 :00 am EST January 27, 1967 

(4 pages). 

(b) Differences existed between the Ground Test Procedures and the In-Flight Check Lists. 

Determination 

Seither the revision nor the differences contributed to the accident. The late 

issuance of the revision, however, prevented test personnel from becoming adequately familiar with the test procedure prior to its use. 

Recommendations 

(a) Test Procedures and Pilot's Checklists that represent the actual Command 

Module configuration be published in final form and reviewed early enough to 

permit adequate preparation and participation of all test organization. 

(b) Timely distribution of test procedures and major changes be made a con- straint to the beginning of any test. 8. FINDIKG 

The fire in Command Module 012 was subsequently simulated closely by a test 

fire in a full-scale mock-up. 

Determination 

risks in flight-configured spacecraft. 

Recommendation 

Full-scale mock-up fire tests can be used to give a realistic appraisal of fire 

Full-scale mock-ups in flight configuration be tested to determine the risk of fire. 

9. FINDING 

The Command Module Environmental Control System design provides a pure 

oxygen atmosphere. 

Determination 

combustibles in the Command Module are not restricted and controlled. 

This atmosphere presents severe fire hazards if the amount and location of 
Recommendations 

(a) The fire safety of the reconfigured Command Module be established by 

full-scale mock-up tests. 

(b) Studies of the use of a diluent gas be continued with particular reference 

to assessing of problems of gas detection and control and the risk of additional 

operations that would be required in the use of a two gas atmosphere. 

10. FINDING 

Deficiencies existed in Command Module design, workmanship and quality 

controrl such as : (a) Components of the Environmental Control System installed in Command 

llodule 012 had a history of many removals and of technical difficulties including 

regulator failures, line failures and Environmental Control Unit failures. nhe 

design and installation features of the Environmental Control Unit makes removal or repair difficult. 

(b) Coolant leakage at solder joints has been a chronic problem. 

(0) The coolant is both corrosive and combustible. 

( d) Deficiencies in design, manufacture, installation, rework and quality 

(e) R'o vibration test was made of a complete fiight-configured stpacecraft. 

(f) Spacecraft design and operating procedures currently require the dis- 

(9) No design features for fire protection were incorporated. 

These deficiencies created an unnecessarily hazardous condition and their concontrol existed in the electrical wiring. 

connecting of electrical connections while powered. 

Determination 

tinuation \vould imperil any future Apollo operations. 

Recommendations 

(a) An in-depth review of all elements, components and assemblies of the 

En-cironmental Control System be conducted to assure its functional and structural integrity and to minimize its contribution to fire risk. 

(71) Present design of soldered joints in plumbing be modified to increase integrity or the joints be replaced with a more structurally reliable configuration. 

( C) The coolant is both corrosive and combustible. 

(d) Review of specifications be conducted, 3-dimensional jigs be used in 

manufacture of wire bundles and rigid inspection at all stages of wiring design, 

manufacture and installation be enforced. 

(e) Vibration tests be conducted of a flight-configured spacecraft. 

(f) The necessity for electrical connections or disconnections with power on within the crew compartment be eliminated. 

(9) Investigation be made of the most effective means of controlling and ex- tinguishing a spacecraft fire. duxiliary breathing oxygen and crew protection from smoke and toxic fumes be provided. 

11. FINDING 

An examination of operating practices showed the following examples of prob- lem areas : a. The number of the open items at the time of shipment of the Command 

Module 012 was not known. There were 113 significant Engineering Orders not 

accomplished at the time Command Module 012 was delivered to NASA ; 623 Engineering Orders were released subsequent to delivery. Of these, 22 were recent 

releases which were not recorded in configuration records at the time of the 

accident. 

b. Established requirements were not followed with regard to the pre-test con- straints list. The list was not completed and signed by designated contractors and 

SASA personnel prior to the test, even though oral agreement to proceed was reached. c. Formulation of and changes to pre-launch test requirements for the Awllo 

spacecraft program were unresponsive to changing conditions. 

d. Non-certified equipment items were installed in the Command Module at 

time of test. e. Discrepancies existed between NAA and NASA MSC specifications regarding 

inclusion and positioning of flammable materials. 

f. The test specification was released in August 19% and was not updated to 

include accumulated changes from release date to date of the test. 
64 

Determination 

Problems of program management and relationships betmeeii Centers and with 

the contractor have led in some cases to insufficient response to changing program 

requirements. 

Reconm endat ion 

Every effort must be made to insure the maximum clarification and understanding of the responsibilities of all the organizations involved, the objective being a fully coordinated and efficient program. 

Xr. FULTOS. I would like to ask General Hedrick, since you are 

Director of Space Headquarters at USAF, clo you have there an 

Inspector General under the U.S. Air Force who is independent and 

makes independent inspections ? 

General HEDRICK. Yes; ve do. 

Jlr. FTTLTOS. Do J-ou need him in Space in the C.S. Air Force? Is 

he valuable ? 

General HEDRICK. Yes. 

Mr. FTLTOI;. I thank you for trying to get an Inspector General 

set LI~ for SASA Either he is not needed in space in the 1-.S. ,lir 

Force or else he is badly needed in NASA. I do feel we need an outside 

independent inspection system that is reportable to the top management of S-lSA. 

As of now, anyone Tho hangs a lemon on a capsule can only complain 

to a contractor, a subcontractor, a man working for the Manned Sppce 

Flight Center, or the particular Center where this project is being 

developecl, or he is required to report to a program director. of course, 

who v-ants to get along with the job. 

If he is down at the launchsite, he will be holding up the lauiiching 

if he thinks there is something which might slightly go wrong and 

probably woii't. So again I recommend to ShSA a strongly indepenclent Iiispector General setup so that we can do especially the So. 0 

recoininenclation to insure adequate understanding and control of the 

engineering and manufacturig details of these subsystems at the subcontractor and vendor level. 

On reconinienclation KO. 5, it would help. 

On recoinmendation No. 6, it ~voulcl help. 

On recommendation KO. 7, it would help. 

I am, of course, interested in the use of Mylar insulation as a.blanket, 

and also interested in the insulation on wires carrying electrical currents under oxygen conditions. 

Has the manufacturer taken off his list or his catalog or limited 

for these purposes these two materials? What has been the result for 

the general public and general business on the investigation ? Are we 

going to limit Teflon and Mylar insulation blanketing ? 

Mr. CORTRIGHT. If I understand your question right, the combustion 

of the Mylar insulation occurred in a ~7ery unusual circumstance, 

namelyMr. FULTOS. I agree vith that. What protection is there for the 

general public and general business with the new information we have ? 

I believe I \Till answer it : The manufacturer, I understand, has taken 

off its catalog lists for these purposes at least the Teflon. 

R4r. CORTRIGHT. I am sorry, Mr. Fulton, I am not awwe of that. 

Mr. FULTOX. One other point I ~~onld liketo ask about is this : When 

there is a coinbiiiation of circumstances resulting in one warning. it 

seems to me incredible that there is not an alternative system that mlght 
65 

turn up a second warning so that the first warning system doesn't 

smother the second. 

Mr. CORTRIGHT.hat is a reasonable observation. The compromise 

always is : how complex can you make the system ? The more modes of 

failure the system can handle, the more complex the system gets. 

This particular system has certain situations of the type you described. We have asked that they be reexamined to see if anything can 

be done about it. It may not be practical to do so. 

By the way, Mr. Chairman, I would like to welcome the last missing 

member of our Board, Mr. Neil Armstrong, who has now arrived. He 

is familiar with one alarm overriding another. He may comment on 

that question. 

Mr. FULTOS. We think that astronauts had better not be cross-examined too closely. We would rather have you fellows respond, and while 

they are orbiting the White House and the Capital now more than the 

moon, we nevertheless give them a littIe immunity, which I think we 

owe them. 

The point I liave always made, and made especially on Apollo 204, 

there is no failure on tlie part of the astronauts in handling the equipment, on running the mission, on the decision to take certain rescue 

operations and the return. The astronauts all performed well n-ithout 

any negligence or failure whatever. 

Is that correct? 

JIr. CORTRIGHT. I guess I donY know of any. 

Mr. FULTOS. How about the Administrator? 

Dr. PAISE. I certainly concur in that statement. 

Xr. FULTOS. How about Air. Arinstrong? 

Mr. ARXSTRONG. I would have to say that there were a number of 

options available to the crev and they didn't inr-estignte every option, 

which in hindsight could have been investigated, but there isnY any 

reason to believe they should have with the information they had 

available to them either. 

Mr. FULTOS. How about General Hedrick ? 

General HEDRICK. I think they perf ormed admirably. Mr. FULTOS. Thank you. 

Mr. HECHLER. Technically, I think this is an outstanding report and 

I like its forthrightness. We can call it the Forthright Cortright 

Reaort. 

Seriously, the recommendations are almost entirely technical in 

nature with the possible exception of parts of KO. 6. In any organization like NASA where you have individuals of high technical competence planning for a very hazardous mission, there has to be niutual 

respect and confidence on the part of those that are using tlie equipment that everything will go right. whereas Murphy's law occasionally crops up. So what really concerns me about both the 1967 fire and 

this accident is that although we hare devised recoininendations which 

take care of correcting tlie tecliiiical aspects. we have done little to 

correct adininistrative deficiencies. 

You really need some critical people who niay not be very popular 

in NASA, they may not get malip invitations to social events, but 

there are people n-ho hal-e a criticnl, skeptical bent in their questions 

about whether or not the contractor has produced safe equipment. 

They must ask tlie kind of questions like. what about the hazards of 
66 

all-oxygen environment, questions like we by hindsight asked in 1967, 

why couldn't yon open a hatch from the inside a little quicker during 

the test. 

I think you need a group of people with this type of inquiring, 

critical mind, that can ask these questions consistently and continuously as the equipment comes from the contractor, to not only watch 

the development and review the procedures according to the manual, 

but to find out if there were any unusual events like the dropping of 

2 inches onto the cement floor and what effect this had. 

I would like to ask Mr. Armstrong if one of the astronauts who has 

made a flight could be placed at the head of the team who could independently ask the kind of questions that the ordinary experts within 

NASA do not ask because they have confidence that even-thing mill 

go right 1 

Mr. CORTRIGHT. If I can interrupt before Mr. Armstrong can answer that question, I would like to correct the error of dropping onto 

the cement floor whicli appeared in the newspaper. This so-called 

drop incident was not like that at all. The tank was assembled in a 

shelf, the shelf was being lifted out of the bay KO. 4 and one bolt 

had not been removed and as a result, the lifting device broke and 

the shelf containing the tn-o oxygen tanks dropped 2 inches back 

onto its mounting brackets. 

Mr. HECHLER. I am glad we got that correction. 

Mr. ARRXSTROSG. I am sure that astronauts who really spend very 

little time in space compared to the amount of time they spend asking 

questions in the course of their job could do such a job as Air. Hechler 

suggested, and we find many other individuals within our Agency 

and without, who are also very penetrating in their inspections and 

could also do such jobs. 

Mr. HECHLER. Would it take someone outside of the Agency coming 

in or could it be done by someone who is necessarily an expert and 

maybe going around with the wheels and have the confidence in the 

equipment which results from just being an expert? 

Mr. ARXSTROXG. I should think there is always some advantage to 

people who are put in this position of having some independent 

authority. 

Blr. HECHLER. I want to ask Dr. Paine if he had any further comment 

on this. 

Dr. PAINE. Of course, Mr. Hechler, we have the Aerospace Safety 

Advisory Panel which is specifically designed to report directly to me 

outside of any other channel. It includes people outside of XA.Sd ~ho 

sit in and re\-iew the procedures we are using. They are penetrating. 

VTe need-not any one magic solution-but we need to take a number 

of different approaches. 

In order to penetrate a system as complex as Apollo to the tremendous depths in which it must be penetrated, and I think we hare a 

beautiful example in this very small thermal switch which was certainly one of the major contributors, it is necessary to have a very large 

organization working on a full-time basis with no other respollsibilities 

snch as our hpollo management system. 

In addition to that, we do need outside people to come in a@ ask 

the very different overall kinds of question, whether or not we Indeed 

have got this set up properly, whether or not the channels of report- 
67 

ing are correct, whether or not we are indeed using the best and most 

modern techniques to attempt to have the entire systeni ferret out such 

questions. 

What we have here before us today is an example of a breakdown of 

a system in which n-e failed with the kinds of gates that n-e have asseinbled to prevent these things froin going through. We failed to detect 

the fact that in the change froin a 28-volt to a 65-~-olt GSE power supply, we failed to test the switch specification. We then have failures 

on the part of additional people later to see to it that this system got 

an adequate test which would expose the fact that this switch, which 

is never called upon to operate in flight, under the ground conditions 

we encountered would fail to operate successfully. 

We have a number of such failures. I haven't had an opportunity to 

go through the report in detail-which I will do-which calls for us 

to reexamine the systems n-e have on place and ask ourselves in detail 

what must be done to make such that in the future we catch things of 

this nature. 

The fact that we did this special detanking procedure on this tank 

Tyliich had never been clone on a previous Apollo mission indicates 

why it was that on Apollo 13 v-e encountered this difficulty vlien we 

had successfully flon-n all the previous Apollo missions. In no case 

had this STT-itcli ever had the opportunity to operate. It vas the special 

cletankiiig proceeding. 

The lesson that we have got to examine here is how it could be that 

we woulcl indeed carry out this special detanking procedure in Cape 

I<ennedy--\vhen we ran into difficulties in detankiiig this tank cluring 

the test period, how it would be that n-e would carry out the procedure 

and not fully examine all the consequences of this. 

There are many questions in the administration end which we must 

reexamine as a follov-up to the job that Air. Cortright and his team 

 ha^-e so ably done. 

Jlr. HECHLER. Thank you, Dr. Paine. 

Mr. KARTH. Thank you, doctor. 

The Chair recognizes Mr. Jiosher. 

Mr. MOSHER. Are you saying that NASA% as yet has not precisely 

identified the point of procedure or the persons in the procedures who 

should hare asked the right questions about the effect of the special 

testing on the pad lvhich fused the thermal switch? 

You haven't yet precisely identified the point or the person where 

the crucial question should ha~e been raised? 

Dr. PAISE. Mr. Mosher, it is my guess that we will never identify 

one particular person that might be called the villain of Apollo 13. 

There are many different failures that hare come to light. 

There was the failure in the switch area. There TT~S the failure in 

pssembling the fill system, which then, in turn. led to the necessity 

at the Cape for the special detanking procedures. 

There were a number of different events which happened along 

the line. Each one of these was necessary. 

Jlr. MOSHER. I wasnY looking for the villain of the piece. I was 

raising essentially the same question that 31r. Hechler and Mr. Fulton 

raised-you haven't precisely identified the person in the future who 

is going to ask these embarrassing or these crucial qnestions? 
68 

Dr. PAINE. That is right. We have not yet made our decision as to 

what changes are necessary in order to preclude such a thing in the 

future. 

Mr. MOSHER. Are there any aspects of this, or any event, not yet 

identified? Is there any remaining mystery as to what happened still 

unexplained ? 

Mr. CORTRIGHT. I guess it is pretty dangerous to say “No” to that 

question. But, at the moment, we don? know of any remaining 

mysteries. 

There was one test which didn‘t turn out quite the m-ay me thought 

it would turn out. It is being rerun at Beech. That mas a fiill-scale 

duplication of the detanking that took place at the Cape in all 

respects. 

When that took place, the switches failed in a different manner than 

they did in our test setup and, as a result, one switch remained closed 

and one open by srirtue of the fact that the terminals melted and 

fell out. 

Mr. MOSHER. So you will still be doing some work? 

Mr. CORTRIGHT. That is right. 

One heater stayed on and one did not ; as a result, the temperatures 

didn’t pet as high as it did in the L4pollo 13, and so the insulation 

wasn’t damaged, although me had done other tests-there are other 

details that need to be cleaned out. 

We pointed out in our letter of transniittal that we plan to reconvene 

a little later in the year to look over any additional tests and analyses 

to see if what we have said here still stands up. 

Mr. MOSHER. You have made several reconmiendations that will take 

time. What about the impact of this on Apo110 147 How much postponement is there going to be? 

Mr. CORTRIGHT. I don‘t knov that there will be any. The recommendations we hare made are generally cast in a two-level type review, 

for example, where we ask the subsystems to be rer-iewed, we are first 

essentially asking for a screening to identify those that we are not SO 

much on top of. It is our feel it can be concluded before A4p~ll~ 14. 

illr. XOSHER. So December is still a good time ? 

Mr. CORTRIGHT, Yes. I think it will be a hard point to meet from the 

changes in Ilardvare that dl be selected. Whether it is possible or 

not, I am not qualified to say. 

Mr. MOSHER. Thank you. 

Blr. DOWSISG. I would like to congratulate Mr. Cortright and the 

board for what I think is an excellent report and a practical one. It 

reminds me of the one we had several years ago. We have complete 

confidence in it. 

Was this the first time that the fan in oxygen tank NO. 2 was 

turned on? 

hfr. CORTRIGHT. KO? sir; the fans and heaters are used \Thenever the 

tank is filled with cryogenic oxygen. They are not used continuously. 

Mr. DOWNING. During the flight ? 

Mr. CORTRIGHT. No. Pardon me. The fans had been turned on several 

times before during the flight. 

Jfr. DOWKISG. Were there other tanks on board which had the 

same switches and thermostats and which did operate properly durillg 

the flight? 
69 

Mr. CORTRIGHT. Oxygen tank No. 1 is essentially identical and it 

operated properly. The hydrogen tanks are similar and they operated 

properly. 

Actually, as I point out in the board report, these particular tanks 

accumulated nearly 3,000 hours of space flight without significant 

problems. 

Mr. DOWNING. They had not been redesigned with the 65-volt 

switch Z 

Mr. CORTRIGHT. No. They have used the 65 volt at the Cape for 

checkout of all of these tanks for pressurization, but not under the 

circumstances of this detanking procedure. 

Let me make sure that is clear. I am not sure that I did this. The difference is that when the heaters are left on during detanking, they are 

running n-hen the tank is almost empty and you don‘t have that large 

quantity of very cold oxygen to keep things cool, so at this point they 

get very hot. That had never happened before. 

Mr. DOWNING. If I read the time chart correctly, there was something more than a minute from the time the fan turned on until 

tlie explosion occurred. Is there anything that the crew could have 

done, in hindsight, or that the ground crew could have done? 

Mr. CORTRIGHT. No, sir; the only thing that could have been done 

v-as to observe the increase in pressure and reduce the troubleshooting time afterward to identify why it happened, but there was nothMr. DOWNING. You termed this a near disaster, which it was. What 

could have happened? 

What did you fear the most? 

Mr. CORTRIGHT. I think that in space it might have been possible 

to rupture a propellant tank in an adjacent ba . It might have failed 

to make the transition that they did to the Lem lifeboat mode. It 

might have occurred at a different point in tlie mission when recovery 

would not have been possible. 

Mr. DOWNING. Was it more of an explosion than an implosion? 

Mr. CORTRIGHT. Yes, sir; I think it is most easily understood as a 

failure in the pressure vessel or its high-pressure tubing due to overheating, a rupture, if you mill, through which high-pressure oxygen 

bursts or streams very rapidly. 

ing that could have been done to save the mission. 

oxygen tank So. 1 more rapidly, not giving tie i” crew adequate time 

Mr. DOWNING. Thank you very much. 

Ah. KARTH. Mr. Winn. 

Mr. WINN. Thank you. 

The review board has done an excellent job in \Thich I concur with 

the remarks of tlie other members of tlie committee. 

I n-ould like to follow up the thought that Mr. Mosher pursued. 

Did anything else in your various tests that you ran give you great 

concern, other than the additional switch, when you were really putting some of these pieces to extreme tests which were shown in the 

movie? Did anything else show up that really bothered you? 

Mr. CORTRIGHT. Yes, sir; these are all spelled out in the board repork 

We were concerned with certain aspects of the basic tank design 

which indicated to us that this ignition might have taken place with 

a tank with good switches in it, in the event tlie insulation would be 

damaged in tlie assembling. 
70 

Until we found the switches, we concentrated J-ery hard on the 

manner in which wiring insulation could be damaged aiid convinced 

ourselves to the point it is still in the report that yes, this could happen with this type of tank design. 

We also were concerned with the amount of potentially combustible 

material in close proximity to electrical sources which could become 

ignition sources. We recommended that something be changed there. 

There was a battery problem on the lunar module which was not 

related to this accident, but it occurred on the way back, and that has 

to be run down. So I think it is not just as simple as this thermostatic 

switch. 

Mr. WIXN. That is what I gathered that you were saying in your 

recommendations, which looked to me as if they were very thorough, 

and you made a statement on page 841, "Where appropriate, X-iSAdesigned standards should be developed." 

In part of your recommendations you say that the review board will 

be called together again shortly. Did I understaiid you to say that? 

Mr. CORTRIGHT.ITe plan one more session ourselves. We are at the 

disposal of the Administrator to reconvene any time he thinks he 

needs us. 

Mr. WIXS. If you haven't developed a program yet, Dr. Paine, TT~O 

in K'aSA, is going to follom- through on these recommendations and if 

additional recommendations are to be made, I don't see how yon are 

going to be able to keep the time schedule for Apollo 14 Then everything is still up in the air. 

Dr. PAINE. This will be examined. After every Apollo mission, a 

great deal of attention is given to going back over all the anomalies 

that have happened. In each mission there have been certain things 

that were unexplained, which had to be dug into, and Mr. Cortright 

has mentioned several additional ones in Apollo 13. 

In no case do we ever fly a mision until we have cleaned up all 

the things to our satisfaction which we have been shown in previous 

flights. 

Mr. WINN. If new parts are needed and new parts hare to be 

designed, built, and tested, I suppose in that case it iTould depend 

on what it is and how important a part it plays in the overall production. When we get down to little m-ires aiid switches, it looks as 

if everything is just as important as the things we hear about. 

Dr. PAIXE. The smallest component is just as important as the 

largest, and ve have just had a very dramatic demonstration of that. 

1 can assure you \re will not fly Apollo 14 until we are satisfied that 

we have fixed up everything that has come to light. 

Mr. WIXN. Thank you very much. 

Mr. KARTH. Mr. Goldwater? 

Mr. GOLDWATER. Why was it necessary to detank this particular 

vehicle ? 

Mr. CORTRIGHT. The procedures at the Cape require that when the 

countdown demonstratmion test is complete, that the tanks be emptied 

and then filled again prior to launch at a later date. I think this is 

partly for safety reasons as a matter of fact. 

Mr. GOLDWATER. This was done on 11 and 122 

Mr. CORTRIGHT. Itis done always. In this particular case, the tank 

would not expel its oxygen in a normal fashion. The way that is done 
71 

is to take the vent line and pressurize the inside of the tank through 

that vent line and that pushes down on the oxygen which pushes 

it up through tlie fill line and out the fill tube. 

li%en there is a loose connection at the top, the grtses you are using 

to pressurize the tank would go in one line and out the other and 

don't pump fluid out with them. That is the problem that was run 

into. 

Mr. GOLDWATER. This happened cluring detanking ? 

Xr. CORTRIGHT. Yes. 

JIr. GOLDWATER. Kotliing \$-as done about it ? 

a problem. 

JIr. GOLDWATER. I see. 

Jlr. CORTRIGHT. Itwas not kno~rii that tlie v-ires had been damaged 

and that the heaters stayed oii continuously, as I told you in the outline of what happened. That was not recognized before Iaunch. 

Mr. GOLDWATER. Did you feel that cluring this detanking period that 

when the temperatures built up, they burned the wires ? 

Mr. GOLDWATER. Could you clarify this cliaiige in the provision froin 

the 28 to the 65 volts power switch specification-why this was important ? 

Jfr. CORTRIGHT. Yes. The spacecraft flies on 28 rolts and Korth 

American Rockwell uses 28 volts. The Kennedy Cape Center uses 65 

1-olts d.c., Beech uses 6.5 volts ax. kt tlie Cape they hare a 65 cl.c., volt 

system. The higher voltages or currents are used to accelerate the tank 

pressurization. When you first fill the tank at low temperature, then to 

build up the pressure at the operating range,. you hare to put heat in 

and you can saw several hours Iy accelerating this, and it seems to 

be an acceptable and desirable procedure froin my point of view, provided ererything is protected from the higher voltage power supplies. 

In this case, that \\-as a cllange, back in 1965, but tlie subcontractor, 

Beech, did not change the switches at that time. They left the sir-itches 

in, or essentially tlie wine switches that were in and these were not 

capable of protecting againqt an orerlieat condition, ~~-1iicIi they neJ-er 

slioulcl have encountered in this cletanlding procedure. 

Mr. CORTRIGHT. It was not recognized that the heater oper a t' 1011 was 

Mr. CORTRIGHT. Yes: the heating claiiiagecl Tire insul a t' 1011. 

Mr. GOLDW.~TER. You are running a 28 volt switch on 6.5 volts ? 

Mr. GOI,DW.\TER. You feel that is what melted the contact ? 

JIr. CORTRIGIIT.Ylii!e tlie switch was closed tlie %volt switch will 

take it. If you attempt to break a c1.c. current, it is difficult to do. 

That arc starts and it wants to hang on, stay there, persist. In the 

process, it erodes, melts, and displaces and. in this case, velcls across 

the tKo contact points. 

Mr. GOLDWATER. Even before the liftoff ? 

Mr. CORTRIGHT. Yes. 

Mr. GOLDWATER. Why did it take so long, 56 hours before the explosion took place ? 

Jfr. CORTRIGTIT. TVe will probably never know. 

JIr. GOLDWATER. YOU said yon were going to elaborate on, which I 

don't think you did, the treinendous pressures that were built up during this explosion, it took some 20-some p.s.i., yet 10 psi. through tlie 

center section could blow the corninand nioclule off the top. 

&. CoRTRIGIiT. Yes. 
72 

Could this happen again with some other system failure! 

Mr. CORTRIGHT. Any time the eiitire face of the command module 

is subjected to about 10 pounds per square inch, it will tear loose from 

the service module. 

This was one of the problems we faced in trying to rationalize 

or understand what happened, and current views based on the Langley 

tests and analyses are that he pressure buildup took place rapidly and 

did not have time to build up against the face of the command module. 

so actually you had high pressures in one part of the structure and 

lesser pressures in the other. 

Mr. GOLDW-~TER. If the pressures didn‘t release out the side. it could 

have gone to the top ? 

Jlr. CORTRIGHT. It could have. 

Mr. IIARTH. Mr. Price? 

Mr. PRICE. Thank you, Jlr. Chairman. 

I want to commend you on your effort. You have ceptainly pointed 

up a lot of things that needed to be pointed up. Perhaps such a board 

should look into the operations of every flight as a means of bettering 

our operation. 

Dr. Paine, doesn’t this point up the need for a rescue system, or 

the thing ve have been talking about, the follom-iiig-on of the shuttle 

and a space station? Had we had such a system in space, there was a 

possibility with the correct modifications that they could har-e attached to a space station and saved their lives? 

Dr. PAINE. This particular accident. and the manner of its occurrence 205,000 miles out on the mission, probably would not hare been 

affected by the capability to launch a rescue mission as we look at it. 

On the other hand, had the accident occurred at another part of 

the mission or in another maiiiier, it is certainly possible that the existence of a space shuttle system or a space station system might hare 

been able to provide some assistance. 

I think it is correct to say that when such systems are available, 

we will all feel a good deal easier about flying men in space. 

You have to recognize we are still in the early days of the Space 

Age, and at the present time we are flying missions with pioneers 

out to explore these new areas, and we do not hare a rescue capability 

for most arts of the mission, particularly, of course, including the 

lunar sur i? ace activities. 

Mr. PRICE. Mr. Cortright, in your first paragraph of your closing 

remarks, you said something about an unforgiving design. Could you 

elaborate on that a little bit ? Who is responsible for an unforgiving 

design ? 

Mr. CORTRIGHT. The prime thought we had in mind, in using that 

word, was the presence of sufficient combustibles in the tank to support 

a rather strenuous fire in there and the combustion paths which permitted this burning 40 get to the vicinity of thin walled, high-pressure 

metal. 

Mr. PRICE. IsNASA responsible for the design as they pass it 011 

for bidding? Is it a factor in the specifications? 

Mr. CORTRIGHT. The process was to have a competitioll in which a 

number of contractors bid and proposed their design. A particular 

subcontractor won. The competition was conducted by the prime contractor and NASS had ail overview responsibility on all of it. The 

ultimate responsibility for accepting the design approacli is KASB’s. 
73 

Dr. PAINE. Icertainly would like to emphasize that, the ultimate 

responsibility for the safety of all our missions is NASA's and we 

fully accept it. 

Mr. PRICE, Neil, I notice on page 2-39 of tlie report, the finding: 

"The crew maneuvered tlie spacecraft to the wrong LM roll attitude 

in preparation for LA1 jettison. This attitude put the CM very close 

to gimbal1 lock which, had it occurred, vould have lost the inertial 

attitude reference essential for an automatic guidance system control 

of reentry." Was this sent up from ground control or could you explain it? 

Mr. ARMSTRONG. No, sir, Mr. Price. It was not bad information 

on the part of the ground. It, in this case, was an error on the part 

of the crew. However, I suspect I might have been guilty of making 

that same type of error, since it was an attitude control situation with 

which they were not familiar as a crew-. 

It was one that was improvised during flight, and there is a certain 

amount of learning involved in this particular control method and 

the interpretation of the clisplays, and they just made a mistake. 

Mr. PRICE. Also, in the testinlong here, it speaks of Manned Spacecraft Center engineers "devised and checked out a procedure for using 

the CAI LiOH canisters to achieve carbon dioxide removal." 

Rfr. Low and I hare been doing some deep sea diving and we began 

finding out about carbon dioxide. This was a critical area in not having 

enough air and rebreathing carbon dioxide, was it not? And why cannot provisions be made in the future. subject to such an eventuality, 

and make it a part of the equipment? It might just mean their survival. 

Mr. CORTRIGRT. Both systems, coinmand module and tlie LM, mere 

designed vith sufficient carbon dioxide removal for their own purposes. The particular failure with the LM lifeboat did not receive much 

attention. 

One of our recommendations is that this be examined to see if the 

coiisumables should be handled or planned in a little different manner 

to enhance this lifeboat capability. That is what you are suggesting, 

and we agree. 

Mr. PRICE. So, if something should happen in the future, we should 

have longer life capability in the LM, eren though we don't now have 

tlie capability-in the future we should develop this so that we can 

well give thein a chance of possible rescue. 

Ke:l, \Tould yon have any comments on that ? 

Mr. ARXSTRONG. I agree with our board chairman that such a thing 

is desirable. This situation was a product of the timing. This particular configuration, the so-called LM lifeboat, was not included as a 

design specification. It 11-as not an intent in tlie original design. It 

was something developed after we had the vehicle and said now, if we 

really get into a problem, what we actually could do is use the LM as 

an aid to help 11s iii an emergency situation. 

That being the case, it is understandable that the particular fittings 

and so on were not compatible, and we recognize now that it would 

certainly be an aid to have them so. 

Mr. PRICE. Recommendation 1 (a) states : 

Remove from contact with oxygen all miring and the unsealed motors which 

can potentially short circuit and ignite adjacent materials. 
74 

What are the potentials of these unsealed motors? It would seem to 

me they should be developed to get away from aiiy potential short 

circuit. Shouldn't you really bear down on this area ? 

Mr. CORTRIGHT. Yes, sir; we should. Those motors and the wiring 

are being looked at very hard. 

Mr. PRICE. So. 2 recommendation states : "The modified cryogenic 

storagy system should be subjected to rigorous requalification program, aiid so forth. 

Shouldn't all these systems be subjected to a rigorous requalification 

program throughout the Apollo 1 

JIr. CORTRIGHT. Any system that is changed has to be properly 

requalified, aiid I suppose. in a sense that recoiiiniend a t' ion was unnecessary because all systems go through a qualification program, but we 

put it in for emphasis. 

Jfr. PRICE. In closing, also 011 reconilnelidatioli So. 6. down in the 

middle of the paragraph J-OU state : 

Furthermore, critical decisions iavolring the flightworthiness of subsystem 

should require the presence and full participation of an expert. 

I am amazed that we don't hare that at present. 

Jfr. CORTRIGHT. In the present case. experts were contacted by 

phone, v-hich is done sometimes. and. in this case, it resulted in some 

confusion and misiiiforniation so that people overlooked the potential 

of an overheating damage. aiid the board is speculating that this might 

not have happened if someone who really knev the inside of that tank 

and all its idiosyncracies had been down there in the coiifereiice on 

detanking. 

Mr. PRICE. It ~~ould seem to me it would be advisable for this type 

of man to be there-that kne\T- the interior workings of ei-ery jointif I were flying I would want that. 

That is all. Jfr. Chairman. 

Mr. IIARTH. Mr. Cortriglit, in addition to your objectivity. I am 

sure that Dr. Paiiie chose you to be head of the revien- board because 

of your competence. and. retrospectively, I wonld say probably it 

was the best choice that could be made. 

In your opinion, in view of your competence, what was most respoiisible for the accident-design, manufacturing. tests, or iiiaiiapeiiieiit ! 

Mr. CORTRIGEIT. I don't think I can aiiswer that by selecting one. 

I think it was an unusual combiliation of things that made this accident 

happen. 

Mr. KARTH. Could you grade those 1.2,3.4? 

Mi. CORTRIGHT. I am afraid I could iiot. 

Mr. KARTH. In your list of recolilllielidatiolls, recoiiiiiiendation 

No. 6, let me just reread that first sentence of the paragraph : 

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. 

1sn:t that standard operating procedure ? 

Mr. CORTRIGHT. So, sir : presentation is iiot necessarily required. 

Mr. IIARTH. Don't you tliiiik it ought to be? 

Mr. CORTRIGI-IT. That is what we are suggesting here ; Tes, ve do. 

311.. IC~RTH. It is rather amazing to me that up to this point in time 

that hasn't been stanclard operating procedure. 
75 

So, 7 : "XASA should conduct a thorough reexamination of all of 

its spacecraft, launcli vehicle, aiid ground systeiiis vliich contain highdensity oxygen," et cetera, et cetera. 

Does my iiieinory serve me properly, after the Apollo 204 fire, 

this ~ras essentially a recommendation ~vliicli had been iiiade at that 

time. 

Because the record doesn't show the shaking of a witness' head 

one way or another, let me point out that one of the witnesses indicates the answer to that question is yes: is that right ? 

Mr. CORTRIGHT. This is our general counsel, George Malley, from 

the Laiigley Research Center, who was also counsel to 204. He seems 

to concur that that was tlie case. 

JIr. ILRTH. As a result of this recoiiiiiieiiclatioii, it is obvious 

that that procedure mts not previously followed. Is that correct ? 

Mr. CORTRIGHT. It is correct to say, here we are witli another oxygen fire on our hands after liaring gone back to look the system overyes. tbat is correct. 

Mr. KARTH. TT'as a thorough reexainination of this particular piece 

of equipment made after the Apollo 204 accident ? 

JIr. CORTRIGHT. In viev of myself and most iiienibers of the board, 

it was not a thorough revieu-. There was some review iiiade of this 

tank and the materials vere once again checked against the so-called 

COJIAT standarcl, but I don't beliere it was as penetrating as it should 

have been. 

llr, KARTH. In your judgment are present manageiiieiit procedures 

entirely adequate to preclude similar future occurrences ? 

JIr. CORTRIGHT. Iwouldn't say that with 100-percent confidence. 

We found the procedures themselves, in general, good, but it was 110ssible to get a nonflightvorthy piece of equipment through, even with 

those procedures, so, until we complete our reexaiiiination of 110~ we 

are doing our business on the subsystems, I i~ould not say that wit11 

confidence. 

On the other hancl, I tliiiili the procedures are good aiid the nianageiiient panel as quite co~iipliineiitary in its review of both the 

procedures and the rigor v-ith which people stick to them and sign 

off all the proper forins and clo all the proper things that are supposed to prevent tliis. 

JIr. KARTH. The hour is late and we have already started a quoruiii 

call. I liad a list of questions I wantecl to ask sou. Because of tlie 

press of time. we will not liave an opportunity to do so. Would you 

prepare answers to them ? 

. 

Jf r. CORTRIGIIT. Yes. sir. 

JIr. KARTH. And submit them for inclusion in the record ! 

(The follo~~ing inforiliation is proviclecl for the record :) 

SATIOSAL XEROSAUTICS ASD SPACE ADMIKISTRATION, 

Jlr. CORTRIGIIT. I-eS, Ah. I'kX'th. 

OFFICE OF THE ADJIISISTRATOR, 

TPasliiizgtoiz, D.C., June 30, 1970. 

Hon. JOSEPH E. KARTH, 

House of Representatices, 

TVnshi?zgton. D.C. 

DEAR JIR. KARTH: This is in resnonse to the questions vou submitted to me cluring the hearing held before the- Committee on Science 'and Astronautics on June 17,1970. 
76 

As Dr. Low and I requested, Dr. Charles D. Harrington, Chairman of the 

Aerospace Safety Advisory Panel, submitted the report of the Panel to us on June 25, 1970, in the form of a letter, a copy of which is attached (TAB A), on the procedures and findings of the Review Board. Based on these reports and on extensive discussions at reviews and meetings held since June 25, Dr. Dale 

Myers, Associate Administrator for Manned Space Flight, has formally submitted to me with his endorsement the final recommendations of Dr. Petrone, 

the Apollo Program Director, to prepare for the Apollo 14 mission. These rec- ommendations are embodied in Dr. Petrone’s memorandum to me of June 27, 

1970, a copy of which is also enclosed (TAB B ) . On the basis of these reports and recommendations, Dr. Low and I have ap proved the following actions to implement the recommendations of the Apollo 13 

Review Board and to carry out the steps recommended by Dr. Petrone and Dr. 

Nyers. First, the recommendations of the Apollo 13 Review Board will be implemented before the Apollo 14 mission is approved for launch. This n ill require postpon- ing the launch date to no earlier than JanuarF 31,1971. 

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 throughout SASA the lessons learned in their areas of responsibility. In 

addition, we will take steps to disseminate widely throughout industry and the 

technical community the lessons of Apollo 13 to prevent recurrences in other areas. Third, the Aerospace Safety Research and Data Institute (ASRDI) at the 

SASA Len% Research Center has been directed to conduct additional research on materials compatibility, ignition, and combustion at various G levels, and on the characteristics of supercritical fluids, as recommended by the Apollo 13 

Review Board. 

Fourth, I have requested that the Aerospace Safety Advisory Panel conduct a review of the management processes utilized by NASA in implementing the 

recommendations of the Apollo 13 Review Board and report to me their views no later than the Apollo 14 Flight Readiness Review. I have also asked Mr. 

Cortright to reconvene the Apollo 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 findings, determinations, or recommendations are nec- essary in light of additional evidence which may become available. 

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 Apollo mission, will permit us to launch Apollo 14 to 

the Fra Mauro region of the moon at the January 31, 1971 launch opportunity. This will also move the planned launch date for Apollo 16 several months to July or August 1971, maintaining the six month interval between launches on which our operations in the Apollo program are now based. However, me will not 

launch Apollo 14 or any other flight unless and until we are confident that we have done everything necessary to eliminate the conditions that caused or con- tributed to the problems we encountered on Apollo 13 and are ready in all other 

respects. 

Question. Are the circumstances of the accident sufficientlv well understood at 

this time to proceed on a firm basis with the Apollo 14 fliuht? 

Answer. Yes. Dr. Low and I have now had an opportunity to study the report 

in detail and to review carefully its recommendations. In our view it is an excellent report based on a thorough and objective investigation and highly competent analysis. It clearly pinpoints the causes of the Apollo 13 accident and sets forth a comprehensive set of recommendations to guide our efforts to prevent the occurrence of similar accidents in the future. 

Qitestion. What is yoitr best estimate of the time and cost to recmer from the 

Apollo 13 accident? 

Answer. 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 Apollo mission, will permit us to launch Apollo 14 to 

the Fra Mauro region of the moon at the January 31, 1971 launch opportunity. 
77 

It is too early to present to you our detailed estimates of the costs and budgetary 

impact of the spacecraft modifications and program changes that we are making. 

Our best current estimate is that the modifications and changes related to the 

actions resulting from 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 ApollO 

budget. Qitestion. Do ~JOZC see tlic need for any major changes in your method of 

operation or procerlztrcs bnscd on tlic Apollo 13 accident experience? 

Answer. SASA’: actions in response to the Board’s recommendations will avoid 

those specific things which led or contributed to the Apollo 13 accident ; and the 

reviexvs and research we have undertaken will help ns avoid future potential 

hazards throughout our programs. 

The reviews now underway throughout SASA in response to the Board’s 

recommendations will, in my view, help us to further strengthen the management 

of Apollo and other ShSA programs. Question. To what cxte?rt tcould yoii cxpcct the results of tlic Apollo 13 accident 

to affect other SASA programs such ns Skylab? 

Answer. The broad effects of the Apollo 13 accident on programs such as Skylab 

have not been determined. Time and cost impact on Skylab, for example, will 

depend on results of decisions and actions taken in the Apollo program and the 

reviews now underway. We do not anticipate any serious implications- on Skylab 

at this time, but re will be continually assessing the situation as these actions 

are taken. 

Certain specific effects have already been eraluated and actions taken relative 

to the Skylab Prograni. These include : assuring that the modifications made to 

the A110110 Service Modules to eliminate the Apollo 13 failure inode will be 

incorporated on Skylab to the extent that the designs are similar ; and applying 

the experience. insight and data gained from Apollo 13 to the Failure Mode and 

Effects Analyses and Single Failure Point Analyses being performed on all 

Skylab flight hardware. 

Question. Do yo21 beliem that NASB can carru out its czirrently planned fiscal 

gear 1971 progrants including costs of tl1e Apollo 13 accident within your original 

budget request to the Congress? 

Anstver. As noted above, we now plan to handle the estimated $10 to $15 million of increased costs within our total Apollo budget. 

Questio?r. To tohut extent are other sgstenis in tlre Apollo cehicla and space- crafts liable to a similar seqiience of ecents leading to tlre Apollo 13 accideiet? 

Ansn-er. We have nom instituted a review of all oxidizer systems in all elements of the Apollo system to be sure, in the light of what we have learned in 

Apollo 13, that materials and energy sources are compatible in these systems, 

and modifications will be made where appropriate. For example, the fuel cell 

oxygen supply valve which now has Teflon-insulated wires in high pressure osygen will be redesigned to eiiminate this hazard. 

I am enclosing (TAB C) for your information a statement which I am presenting to the Senate Committee on Aeronautical and Space Sciences at a hearing 

this morning which discusses these actions in greater detail the actions we plan 

to take in responqe to the Board’s recommendations. 

I hare the utmost confidence that the SASA teain can fix the dpollo 13 problem 

and strengthen its operations to minimize the chances of future problems. 

We will keep you and the Committee informed of developments. 

MAKUFACTURING AXD TEST 

Qteestimi. Did the manirfactlrre, qztalificution and testing of th c Service Xodtile 

oxugen system conform to best practices at tlre time of its decelopnient? 

Answer. The design was difficult to manufacture. but good practices were followed to help insure against manufacturing defects. Good testing procedures were followed, but the tests did not include a test of the thermostatic switches functioning under load. 

Question. Were the latest impr0z;enimts ill nianirfactzcre aiid tcst incorporated 

in thc manufacture of the Serrice Xodule Ox,ygen systenz drtring tlie progress of 

tlie progranz? 

Answer. Many improvements were incorporated in the nianufacture of the 

oxygen tanks during the progress of the program. These included the use of special tools, jigs, and fixtures ; improved assembly, cleaning and inspection procedures, and more thorough and improved testing and checkout operations. 

47-591 0-70-6 
78 

Question. Could the problenz of the thermal switches have been anticipated 

and corrected in tlie original testi,ig and nianiifacture of the oxygen tanks? 

Answer. If the design qualification or the flight unit acceptance testing of the 

oxygen tanks had included a functional test of the thermostatic switch interrupting the 65 volt DC, 6 amp ground power load, the potential problem could 

have been uncovered and corrected. 

Questimi. The latinch crews handling the oxygen systenz tests pvior to Zaiinclr 

of Apollo 13 were unaware of tlie potential problem of the tlicriiiul switclies in 

the oxygen tank. Was docunzentation and expert support personnel from industry 

and NASA auailable to diagnose this problent? 

Answer. Although adequate documentation and expert personnel necessary to uncover the potential thermostatic switch problem were not available at KSC, 

they did not exist among USC, Sorth American Rockwell and Beech Aircraft. 

However, the switch probleni probably could not have been readily uncovered as demonstrated by the fact that it took considerable time and effort of uany people 

to uncover after the flight, when it was not just a potential problem. 

XISSION ANALYSIS 

Question. Had the Apollo 13 accident occzrrred in other portions of the flight 

do you beliezje it would haae been possible to have recovered the astronauts? 

Answer. The Board did not review in detail the possible consequences of SJI 

oxygen systeni failure at other times during the mission. Launch pad abort procedures and the launch escape system are designed to cope with emergencies on the 

pad or during the early portion of boost. Obviously, recovery would hare been 

earlier and more simple had the accident occurred in earth orbit. Once the LJIl 

separated from the CSJI in lunar orbit, recovery would have been more difficult, 

and, in some cases, perhaps impossible. However, as pointed out in our testimony, the possibilities of recovery would have depended on the actions Tvhich 

could be taken under the precise circumstances involved. 

Question. Based on the outstanding perforniance of tlfc astronauts, ground coittrollers and supporting personnel do you believe that new OT Chan@ proce- dures, equipnient or techniques should be proljidecl to improve the probability of 

recotmy in the event of an accident? 

Answer. Recommendations 3 and 4 of the Board’s report recommend that the 

Manned Spacecraft Center (JlSC ) consider several specific changes in equipment 

and operating procedures to improve the possibility of recovery in the event of an accident, and that consumables and ecpiipineiit in the LJI ancl CAI be revieIved 

to determine if their potential utility in the “LJl-lifeboat” mode should be en- hanced. Certain tradeoffs must be considered with regard to these recommendations, since the addition of further reduiidancy or complesity might reduce the 

probability of mission success and crew safety. Question. Are we taking adGantage of our new extensive operational experi- ence to assure niaxiniunz safety for the astronnuts both in ternis of surcical 

equipment and procedures ? 

Answer. We learned a great deal froni the Apollo 13 accident regarding the 

ability of the spacecraft, Mission Control ancl the wen- to function under es- tremely adverse conditions. The knowledge gained from this experience ii being 

used to enhance and improve simulation ancl training methods to better prepare 

future crews for dealing with emergencies. and the Board has recommended re- view of equipment and procedures in light of this experience. 

JXANAGEJI EST 

Question. In the sequence of design, mantifacture and test were procedztres for quality assurance and reliability fully coniplied tcith by all lccels of con- tractor and NASA nianagenient? 

Answer. The review of the Board and its Panels of the oxygen tanli system 

indicated that the procedures for quality assurance and reliability were fully complied with. 

Question. Where modifications were required to the Service Jfodztle oxygen 

systenz, was management visibility within SSSA and the contractors sztflicient to understand potential problem areas? 

Answer. Change control procedures were in effect and followed in the course of 

design, manufacture and test of the oxygen tank system. T’isibility was afforded 

to appropriate levels of management during the course of the \vorli. As the Report of the Review Board states. less detailed procedures were in effect in the 
early history of the oxygen tank system than are now in effect. The Board further concluded that attention in the design of the system was primarily devoted 

to its thermodynamic performance, with relatively less attention given to other 

design details. 

Question. Arc the ntanagement procedzires currently in effect suflicient to 

procide SASA and contractor nzanagcnzent adeqnatc infornzutioii fo preeludc 

sinti.7ar occurrences on fittiire flights? 

hnsiyer. The management procedures in effect provide a great deal of inforniation and our review indicated that the procedures were followed. Essentially 

all the information Ivliich the Board used in tracing tlie historj of the oxygen 

tank system was available in the records of XASA or its contractors. We found 

that there are extensive documentation and procedural controls in effect and it was not obvious to us that major additional procedures are necessary to add 

to the information that is available. 

It should be noted, lion-ever, that the Board recomniencled a reassessment of 

subsystems to insure adequate understanding and control of the details of the 

subsystems at the subcontractor and rendor level. The Board also believes that some specific procedural improvenients are warranted and made recommendations OII those points. 

DESIGX 

Question. Wws tlie basic design of tlie oxygen system of the Sercice Vodrtle 

sound in concept? 

Answer. The basic design of the oxFgen srstem of the Service Module is con- sidered sound in concept, and no changes in basic system design have been 

recommended. The detailed design of the interior components of the oxygen tank 

included a number of cleficiencies which are identified in the Board Report and 

Appendices. The design of these components should be modifiec!, and this re- design is undenvay. 

Question. At tlic tinic of design of tlte oxygen system in the Service Vodzile 

in 1965-1966 icere all of the releeant factors of design k?zotcn, at the tinae taken 

into consideration .? 

Answer. The oxygen tank was originally designed in the 1962 to 1963 tinie 

period. This was designated the Blocli 1 system. In 196.5-1966, slight niodifica 

tions were made-primarily to enhance reliability. This modified systeni was 

designated Block 2. The principal change from Block 1 was the provision of 

independent circuits for each of the fan niotors and heater elements, thus providing functional redundancy for each of these motors. 

In general. the relevant factors of design representing the state-of-the-art at 

the time were incorporated in the design. To cite a few examples : 

(1) The material of the pressure yessel is most suitable for this service. 

(2) Storing the oxygen in the supercritical state was appropriate. By maintaining the oxygen in this single phase high density state, withdrawing the 

oxygen for use in simplified. high storage efficiency is obtained. and alo~h during 

acceleration is avoided. 

13) Providing a means for mixing or stirring the fluid was required to assure a homogenous fluid. This avoided the uncertainties associated with the then 

inipedectly understood behavior of fluids under zero-G conditions. 

On tlie other hand. the factor variously ternied mminfacture-ability or proCncibility was not taken into account appropriate1.r. This factor includes such 

considerations as inspectability and testability. It is difficult to install the internal 

coniiionents of the tank srsteni, part of the procedure being “blind.” This process 

ip conducire to \Tire damage that can go undetected without visual inspection. 

Such inspection is not possible xvith this configuration. 

Thus. in this respect it may be said that all of the factors of the design were not taken into account appropriately. 

Question. Did SlSd at thc time of design of the oxygen tank system ham a definite proccditrc for rcpdnting tlie egzripnzent as nezc k?ioKlcdge icas gained 

tliroicglr. operation nnd tests? 

Answer. Yes. the nianagenient procedures in use in Apollo did provide for 

updating designs as required. 

Qitestion. Is it ~icecssavy to conzpletety redesign the Sercice Xodule oxygen 

system or cuii clian,gcs be niwde icliicli will clinzinate potential causes of the 

.lpol7o 13 accidcnt? 

Anwer. So, a complete redesign is not necessary. Changes to the internal 

coniimnents of the oxygen tank and the fuel cell shut-off ralves have been recommended and work is proceeding on these changes. 
80 

Question. Are other oxygen tanks toithiio t71e Bpollo ?;chicle and spacecraft srtbject to tlie sanie problem? 

Answer. Each reiiiaining Service Module presently iiicludes tanks identical to 

oxFgen tank number 2 in .4pollo 13. These will be modified. KO other oxygen 

tanks in the Apollo spacecraft are closely similar to these tanks. !@he Board 

recommended that all high pressure oxygen systeins in the spacecraft be re examined. 

Mr. KARTH. Had the accident occurred at any other time during the 

mission, when would it have been unrecoverable ? 

Mr. CORTRIGHT. After separation froin the lunar module for oneI will ask Mr. Armstrong to answer that question. 

Mr. ARJISTROXG. Ithink in general that answer is probably sufficient as it stands. 

I have found in these kinds of situations that people vhen pressed 

can usually come up with some effective surriral procedures which 

are coinpletely iionstandard and woulcl be unacceptable before the 

fact, but n-lien they are the only last ditch effort, that you find, in fact, 

they will work, and we really don't know hov- long people will live. 

We are talking about i'uiining out of coiisuinable oxygen, coolants, 

and, in order to say how long one miglit live in those conditions, you 

hare to predict physiological factors of individuals aiid n-hen aiid how 

long in a high CO, atmosphere they might exist, we don't have good 

data. 

There iiiiglit be some cases where they miglit survive, but to predict 

their survival would be difficult. I woulcl say his answer as it stood 

froin the point of view of rigor is correct in itself. 

Mr. KARTII. Are there anv further questions ? 

Mr. Fultoii. 

Mr. FULTOS. This brings up the qnestioii that this was actually the 

same equipment that was operative in both Apollo 11 aiid 12, was it 

not z 

Mr. CORTRIGIIT. Yes, sir. 

Mr. FCLTOS. But if something else hacl happened, it would seem 

to me that the equipment x-oiild have operated all right. That ~oulcl 

pretty xell eliininate the equipment as an individual inducing cause 

on present ~vithout soiiietliiiig else having occurred. 

. 

Could I ask Mr. Armstrong to coniinent ? 

Air. ARMSTRONG. Yes, sir; Mr. Fulton. 

As you know, we spend a great deal of our time in the preparation 

for emergencies in our training and in our thought processes, planning 

for these flights. 

In the case of crew members, certainly about 75 percent of their time 

is involved in planning' for these emergency situations, so we are not 

at all surprised when they occiir: as a matter of fact, \ve are probably 

surprised that so few of them occiir in our real flights. 

* 

Mr. FVLTOS. I am ready to go on any trip. Please note. 

Mr. KARTH. 7T'e are ready to send liini too. 

JIr. FULTON. I have had one person recommend a one-way trip. 

Nr, ARMSTRONG. The problem occurs when you have a coinbination 

of circumstances, aiid that is the sitnation which existed here. This 

supersedes our ability to actually, substantially and correctly react and 

predict those kinds of combinations of failure circumstances. 
81 

Mr. FXLTOX. May I coinniend Mr. PsLine, the Administrator, on his 

good comments on the safety panel, aiid may I ask that the accoinplishinelits of the safety panel be put in the record at this point? 

Mr. KARTH. No objection. 

(Infornabion requested for the record follows :) 

The Panel reinforces the continuing attention of SASS and its contractors to 

risk assessment land the formalization of the hazard identification and control 

process. Given the dynaniics of the development process, the multitude of design 

and operational decisions and the broad span of technology inherent in SASA’s 

programs it was recognized that the Panel could assess at most a very limited nuiiiber of these decisions. Therefore it was iiiutuiallp agreed that tlie Panel’s 

effectiveness nonld lie in focusing on the evolution of tlie risk management systeiiis and policies. The Panel’s first year was spent in a surrey of the Apollo program management 

system and the system for hazard identification and risk assessment. This also 

enabled the Panel to assess the impact of agency staff activities. The Panel re- viewed technical management policies and controls at the system level. Attention was focused on configuration management because of the importance of a systeni to define the configuration “as designed” and “as built,” its test history and the waivers and devitations accepted as risks. The Panel was also particularly interested in the institutionalization of systeni safety given the structure of the fundamental risk nianageme_nt system. Because the dpollo prograin was in an advanced stiage when the Panel ~vas established it was difficult to evaluate tlie historical adequacy of the Apollo risk management system. Therefore, 

the Panel monitored the system as it provided an assessment of iiiission risks. 

The Panel gave specific attention to the processes for re-evaluation of possible 

worst case failure inodes and definition of the safety factors in life support systems and consumables. 

The Panel’s review of tlie Apollo program involved staff and program elenients 

at SASA Headquarters, the manned space Right centers, and the majority of 

principal contractors for tlie spacecraft, launch vehicles and Apollo mission 

support. The Panel met in session twenty-two days. While the Panel had not 

studied any area sufficiently to evaluate it in depth, the technical management 

background of the members permitted them to comment selectively on the 

described systems. 

The Panel has recently completed an assessment of the management process for tlie evaluation of risks inherent in reducing Saturn static testing and launch 

operations, as well as a review of the inrestigation process involved in the 

LLTT’/LLRT’ accidents. 

The Panel has also been asked by the Administrator to reviev the hazard 

identification and risk aesessinent system on tlie SERT’h/nuclear stage, the 

space station and space shuttle. Involvement in the definition phases of program 

development promises increasing effectiveness for the Panel as tlie programs 

mature. 

Currently tlie Panel is involved in an assescinent of the procedures, and the 

findings, deterniinations and recoxnniendations of the Apollo 13 Review Board. 

JIr. FL-LTOS. I u-oiild like to coniineiicl Jfr. Runisfeld and those of 

us vho put in the bill the recoiiiiiieiidatioiis of the Safety Panel. 

I would like, along thc lines that have been discussed here. some 

further management inquiry and a report to be iiiade on IIOK we 

can get better inspection procedures so that it is independent, so 

that someboc1;v outside the line of either production, or the manageiiient, or the program director. the Center or the launch area, can 

be appealed to by anyone who feels that sonicthinp should be looked 

into further. 

At the present time. I feel that there is the pressure to pet the job 

done, aiid it is too iiiucli to exiiect any particular indiriclual to step 

clear out of line if he has some ideas. 

If I could make that suggestion, I 1~ould hope that further safety 

procedures will be looked into on the nianapeinent level. 
82 

I think KASh is doing a fine job, and I ail1 pleased this committee 

has met this inorniiig, looking into this excellent report. 

Nr. KARTH. On that note, I think, Dr. Paine, it will be necessary 

for us to conclude the hearings today, aiicl I n-aiit to thank you again. 

Mr. Cortright, and members of the reriew board for preparing this 

report. 

If the chairinan feels it is necessary to go further, I am sure the? 

will be in touch. Thank you. 

(TYhereupon, at 12 :20 pin., the coiiimittee was adjourned.) 
83 

APPENDIX 

REPORT OF 

APOLLO 13 REVIEW BOARD 

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 
84 

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 

APOLLO 13 REVDN BOARD 

June 15, 1970 

The Honorable Thomas 0. Paine 

Administrator 

National Aeronautics and Space Administration 

Washington, D.C. 20546 

Dear DT. Paine: 

Pursuant to your directives of April 17 and April 21, 1970, I am 

transmitting the final Report of the Apollo 13 Review Board. 

Concurrent with this transmittal, I have recessed the Board, subject 

to call. 

We plan to reconvene later this year when most of the remaining 

special tests have been completed, in order to review the results 

of these tests to determine whether any modifications to our 

findings, determinations, or reconanendations are necessary. In 

addition, we will stand ready to reconvene at your request. 

Sincerely yours, 

Edgar M. Cortright 

Chairman 
85 

PREFACE 

The Apolio 13 acclden:, which aborted man's third mission to explore 

the surface of the Lloon, is a harsh reminder of the immense difficulty 

of this mdertaking. 

The total Apollo system of ground complexes, launch vehlcle, and 

spacecraft constitutes ;he nost ambizious and demanding engineering 

development ever undertaken by man. For these rrissicns to succeed, both 

men and equipment must oerform to near perfection. That this system has 

already resulted in two successful 1mar surface explorations is a tribute 

tc those men and women who conceived, designed, built, and flew it. 

Perfection is not only difficult to achieve, but difficult to maintain. The imperfec-ion in A-pollo 13 constituted a near disaster, averted 

only by outstaEdlzg perforxance on ',he part of the crew and the ground 

control tea?? ~.<hLc?. susported then. 

Tke ksollo 13 Fexiiew Joard was charged with the resFonsibilities 

of reviewir& the zirc.;-.stances surrounding the accident, of establishing 

the grobable causes or^ :he eccident, of assessing the effectiveness of 

A-igz: recovery aotiors, of reporting these findings, and of developing 

recor.mer.da3tiors fcr corrective or other actions. The Board has lade 

eveq effort tc carry ou: its assignment in a thoro.agh, objective, and 

ir.rarria1 manner. In doing so, tke Board r.ade effective me of the 

failure analyses and correcive action studies carried out by the ::armed 

Spacecraft Center 2nd was .Jey bpressed with the dedicaticn an6 objectivity of this effor:. 

E?, , 

Yr.e 30~r.d feels that the nature of the Apollo 13 eqGipnect failure 

tolds ::.portart lessons which, vher applied to fLture nissions, will 

coztribute to the safeiy and effecziveness of manred space flight. 
CSM combined t

LM overhead hatch 

Lunar module ascent stage Comrrand rnadule 

t

Lunar module descent stage 

I- 

< 

Apollo 13 space vehicl
87 
88 

.8 Inflight phokcgraph of service module showing 3amg.- ;G uaj- 4. 
89 

TABLE OF COKCEhTS 

Page 

LETT53 OF TMSMITTAL 

PWFACE ............................ iii 

TABLE OF CONTENTS ....................... ix 

CHAPTEB 1 - AUTHORITIES 

Memorandum, April 17, 1970, from Administrator and 

Deputy Administrator to Mi-. Edgar M. Cortright , , . . , . 1-1 

Memorandum, April 21, 1970, from Administrator and 

Deputy Administrator to Mr. Edgar M. Cortright . , , . . , 1-4 

Memorandum, April 20, 1970, from Administrator and 

Deputy Administrator to Dr. Charles D. Harrington, 

Chairman, Aerospace Safety Advisory Panel ......... 1-6 

Memorandum, April 20, 1970, from Administrator to 

Mr. Dale D. hiyers, Associate Administrator for 

Mame3 SFace Plight .................... 1-7 

NASA Management Instruction 8621.1, Subject: 

Mission Failure Investigation Policy and Procedures, 

April 14, 1966 ...................... 1-8 

NASA Management Inswuction 1156.14, Subject: 

Aerospace Safety Advisory Panel, December 7, 1967 ..... 1-10 

C32TER 2 - BOARD HISTORY AID PROCEDURES 

Part 1. Summary of Board History ar.d Procedures ...... 2-1 

Part 2. Biography of Board Members, Observers, and 

Panel Chairmen. .................. 2-3 

Part 3. soar3 Organization and General Assignments 

for Board Panels .................. 2-17 

Part 4. Summary of Board Activities ............ 2-27 

ix 
90 

CHAPTER 3 - DESCRIPTION OF AWLLO 13 SPACE VEHICLE AND EISSION 

Part 1. Apollo/Saturn V Space Vehicle , , . . . , , , . . 

Part 2. Apollo 13 Mission Description , . . , , , , , , , 

CHAPTER 4 - REVIE'd AND ANALYSIS OF APOLLO 13 ACCIDEW 

Part 1. Introduction . . . . . . . . . . . . . . . . . . 

Part 2. Oxygen Tam No. :-istory . . , . . . . . . , . . . 

Part 3. Apollo 13 Flight . . . . . . . . . . . . . . . . 

?art 4. Sunvary Analysis of the Accident . . . . . . , , 

Part 5. Apollo 13 Recovery . . . , . . . . , . . . . . . 

CHAPTER 5 - FINDINGS, DETERMTTUTIONS, AND RECO~ETDATIONS 

Part 1. Introduction . . . . . . . . . . . . . . . . . . 

Part 2. Assessment of Accident . . . . . . . . . . . . . 

Part 3. Supporting Considerations . , . . . . . . . . . . 

Part 4. Recommendstions , , . , , , . , . . . . . . . . . 

Page 

3-2 

3-26 

1-1 

4-2 

4-25 

1-36 

4-14 

5-1 

5-5 

5-12 

5-43 

APPENDIX A - BASELINE DATA: APOLLO 13 FLIGHT SYSTEM AND 3PE>UTIONS 

APPENDIX B - REPORT 3F MISSION EVEUTS PANEL 

APPENDIX C - REFQRT OF M4NDiFACTURIN'; AND TEST PANEL 

APPENDIX D - REP3RT OF DESIGN PANEL 

APENDIX E - REPORT OF PROJECT M4N4GWNT PAmL 

APPZYDIX F - SPECIAL TESTS AND ANALYSES 

APPZNDIX G - BOAR3 ADMNISTRATIVE PRODECURES 

AP?EXDIX H - B3.QD RE;E4SES AND PRESS STATEMEWE 

X 
91 

CHAPTER 1 

AUTHORITIES 
92 

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 

WASHINGTON. D.C. 20546 

OFFICE OF THE ADMINISTRATOR April 17, 1970 

TO : Mr. Edgar M. Cortright 

SUBJECT : Establishment of Apollo 13 Review Board 

REFERENCES: (a) NMI 8621.1 - Mission Failure Investigation Policy 

and Procedures 

(b) NMI 1156.14 - Aerospace Safety Advisory Panel 

1. It is NASA policy as stated in Reference (a) "to investigate and 

document the causes of all major mission failures which occur in the 

conduct of it5 space and aeronautical activities and to take appropriate 

corrective actions as a result of the findings and recorrmendations." 

2. 

craft which jeopardized human life and caused failure of the Apollo 13 

lunar mission, we hereby 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 NASA and other Government agencies. After consultation with you, we will: 

Because of the serious nature of the accident of the Apollo 13 space- 

(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 NASA Office of Public Affairs, to the public. The Board will report 

its findings and recommendations directly to US. 

3. The Board will: 

(a) Review the circumstances surrounding the accident to the spacecraft which occurred during the flight of Apollo 13 and the subsequent 

flight and ground actions taken to recover, 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. 

1-1 
(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 Aerospace 

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 organization, 

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 Administrator for 

Manned Space Flight to: 

(a) Assure that all elements of the Office 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 reviews, 

studies, and supporting actions as are required to develop recomendations to us on corrective measures to be taken prior to the Apollo 14 

mission with respect to hardware, operational procedures, and other 

aspects of the Apollo program. 

1-2 

47-591 0 - 70 - 7 
94 

7. All elements of NASA will cooperate with the Board and provide full 

suuport .. within their areas of responsibility. 

Georee -- M. Low 

Deputy Adminis t ra t or T. 0. Paine 

Administrator 

1- 3 
95 

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 

WASHINGTON, D.C. 20546 

OFFICE OF THE ADMINISTRATOR April 21, 1970 

TO : Mr. Edgar M. Cortright 

SUBJECT : Membership of Apollo 13 Review Board 

Reference: Memorandum to you of April 17, subject: Establishment of 

Apollo 13 Review Board 

In accordance with paragraph 2(a) of Reference (a), the membership of 

the Apollo 13 Review Board is established as follows: 

Members : 

Mr. Edgar M. Cortright, Chairman (Director, Langley Research Center) 

Mr. Robert F. Allnutt (Assistant to the Administrator, NASA Hqs.) 

Mr. Neil Armstrong (Astronaut, Manned Spacecraft Center) 

Dr. John F. Clark (Director, Goddard Space Flight Center) 

Brig. General Walter R. Hedrick, Jr. (Director of Space, DCS/R&D, 

Mr. Vincent L. Johnson (Deputy Associate Administrator-Engineering, 

Mr. Milton Klein (Manager, AEC-NASA Space Nuclear Propulsion Office) 

Dr. Hans M. Mark (Director, Ames Research Center) 

Hqs., USAF) 

Office of Space Science and Applications) 

Counsel : 

Mr. George Malley (Chief Counsel, Langley Research Center) 

OMSF Technical Support: 

Mr. Charles W. Mathews (Deputy Associate Administrator, Office of 

Manned Space Flight) 

Observers: 

Mr. William A. Anders (Executive Secretary, National Aeronautics 

and Space Council) 
96 

Dr. Charles D. Harrington (Chairman, NASA Aerospace Safety 

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

Advisory Panel) 

Data Institute, Lewis Research Center) 

Congressional Liaison: 

Mr. Gerald 3. Mossinghoff (Office of Legislative Affairs, NASA Hqs.) 

Public Affairs Liaison: 

Mr. Brian Duff (Public Affairs Officer, Manned Spacecraft Center) 

In accordance with applicable NASA instruction, you are authorized to * 

appoint such experts and additional consultants as are required for 

the effective operations of the Board. 

George M. Low 

Deputy Administrator 

T. 0. Paine 

Administrator 

1- 5 
97 

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 

WASHINGTON, D C 20546 

April 20, 1970 

OFFICE OF THE ADMINISTRATOR 

TO : Dr. Charles D. Harrington 

Chairman, Aerospace Safety Advisory Panel 

SUBJECT : Review of Procedures and Findings of Apollo 13 Review Board 

Attachment: (a) Memorandum dated April 17, 1970, 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 1156.14 - Aerospace Safety Advisory Panel 

1. In accordance with References (a) and (b), the Aerospace Safety 

Advisory Panel (hereafter referred to as the Panel) is requested to 

review the procedures and findings of the Apollo 13 Review Board (hereafter referred to as the Board) established by Attachment (a). 

2. The procedures established by the Board will be made available to the 

Panel for review and comment as provided in paragraph 4(a) of Attachment (a). 

3. As Chairman of che Panel, you are designated an Observer on the Board. 

In this capacity, you, or another member of the Panel designated by you, 

are authorized to be present at those regular meetings of the Board you 

desire to attend. You are also authorized to receive oral progress reports from the Chairman 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. 

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. 

The final report and any interim reports of the Board will be made 

George M. Low 

Deputy Administrator 

OL “PT. 0. Paine 

Administrator 

Enclosure 

cc: Mr. Edgar M. Cortright, Chairman, Apollo 13 Review Board 

M/Mr. Dale Myers 

1-6 
98 

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 

WASHINGTON. D C 20546 

April 20, 1970 OFFICE or THE ADWNISTRAIOII 

TO : MK. Dale D. Myers 

Associate Administrator for Manned Space Flight 

SUBJECT : Apollo 13 Review 

References: (a) Memorandum dated April 17, 1970, to Mr. Edgar M. 

Cortright, subject: Establishment of Apollo 13 

Review Board 

(b) Memorandum dated April 20, 1970, to Dr. Charles 

D. Harrington, subject: Review of Procedures 

and Findings of Apollo 13 Review Board 

1. As indicated in paragraph 6 of Reference (a), you are directed to: 

Assure that all elements of the Office of Manned Space 

Flight cooperate fully with the Board in providing records, 

data, and technical support as requested. 

Undertake through the regular OMSF organization such reviews, 

studies, and supporting actions as are required to develop 

timely recommendations to us on corrective measures to be 

taken prior to the Apollo 14 mission with respect to hardware, operational procedures, flight crews, and other aspects 

of the Apollo program. 

2. The recommendations referred to in paragraph l(b) above should be 

submitted to us in such form and at such time as you deem appropriate, 

but a report should be submitted no later than ten days after the 

Apollo 13 Review Board submits its final report. 

3. The assignments to the Apollo 13 Review Board and to the AerOSpace Safety Advisory Panel by References (a) and (b), respectively, 

in no way relieve you of your continuing full responsibility for the 

conduct of the Apollo and other OMSF programs. 

443 <=r;b wu 

Deputy Administrator Administrator 

cc: MK. Edgar M. Cortright, Chairman, Apollo 13 Review Board 

MK. Charles D. Harrington, Chairman, Aerospace Safety Advisory Panel 
99 

April 14, 1966 

Hiedive dde 

Management Instruction 

SUBJECT: MISSION FAILURE INVESTIGATION POLICY AND PROCEDURES 

1. 

2. 

3. 

4. 

5. 

PURPOSE 

This Instruction establishes the policy and procedures for investigating 

and doclrmenting the causes of all major mission failures which occur in the 

conduct of NASA space and aeronautical activities. 

APPLICABILITY 

This Instruction is applicable to NASA Headquarters and field installations, 

DEFINITION 

For Lhe purpose of :his Instruction, the following term shall apply: 

In general, a failure is defined as not achieving a major mission 

objective . 

PoLIcy 

a, It is NASA policy to investigate and document 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. 

b. The Deputy Administrator may conduct independent investigations 

of major failures in addition to those investigations required of 

the Officials-in-Charge of Headquarters Program Offices as set 

forth in paragraph 5a. 

PROCEDURES 

a. Officials-in-Charge of Headquarters Program Offices are responsible, 

within their assigned areas, for: 

(1) Informing promptly the Deputy Administrator of each major 

failure and apprising him of the nature of the failure, status 

of investigations, and corrective or other actions which are 

or will be taken. 

1- 8 
NMI 8621.1 

(2) Determining the 

100 

auses or probable uses 

taking corrective or other actions, and su 

April 14, 1966 

E all failures, 

nitting written 

reports of such determinations and-actions to the Deputy 

Administrator. 

b. When the Deputy Administrator decides to conduct an independent 

investigation, he will: 

(1) Establish a (name of project) Review Board, comprised of appropriate NASA officials; 

(2) Define the specific responsibilities of each Board, encompassing 

such tasks as: 

Reviewing the findings, determinations and corrective or 

other actions which have been developed by contractors, 

field installations and the Official-in-Charge of cognizant 

Headquarters Program Office and presenting the Board's 

conclusions as to their adequacy to the Deputy Administrator. 

Reviewing the findings during the course of investigations 

with cognizant field installation and Headquarters officials. 

Reconunending such additional steps (for example additional 

tests) as are considered desirable, to determine the technical and operational causes or probable causes of failure, 

and to obtain evidence of nontechpical contributing factors. 

0 

(d) Developing recommendations for corrective and other actions, 

based 011 all information available to the Board. 

(e) Docui~n~ing findicgs, determinations and recommendations 

for corrective or other actions and submitting such docimentation to the Qeputy Administrator. 

C. Procedures for implementing the Board's recommendations shall be 

determined by the Deputy Administrator. 

6. CANCELLATION 

NASA Management Manual Instruction 4-1-7 (T.S. 760), March 26 1964. 

Deputy Administrator 

DISTRIBUTION : 

SDL 1 

1- 9 
101 

Mectivr dale 

Management Instruction 

SUBJECT: AEROSPACE SAFETY ADVISORY PANEL 

1. PURPOSE 

This Instruction sets forth the authority for, and The 

duties , procedures, organiz&tion, 2nd su2port of the 

Aerospace Safety Advisory Panel. 

2. AUTHORITY 

The Aerospace Safety Advisory ?me1 (hereafter called the 

"Panel") was established under Section 6 of the National 

Aeronautics anci Space Administration Authorization Act, 

1963 (PL 90-67, 50th Congress, 81 Stat. 168, 170). Since 

the Panel was established by statute, its foriiiation an6 use are not subject to the provisions of Executive Or>tier 

11007 or of IdMI 1150.2, except to the extent that such 

provisions are made applicable to the Panel under this 

Instruction. 

3. DUTIES 

a. The duties of the Panel are set forth in Section 6 

of the ?:ational Aeronautics and Space Adninistration 

buthorization Act, 1368, as follows: 

"The Panel shall review safety stuciies. 2nd 

operations plans referred to it and shall 

inake recorts thereon, shall advise the 

Administrator with respect to the hazards 

of proposed or existing facilities and proposed operations and with respect to the 

adequacy of proposed or existing safety 

standards, and shall perform such other 

duties as the Administrator may request ,I' 

b. Pursuant to carrying out its statutory duties, the 

Panel will review, evaluate, and advise on all 

elements of iJASA's safety system, including 

especially the industrial safety, systems safety, 

1-10 
NMI 1156.14 

102 

December 7, 1967 

and public safety activities, and the nianagement of 

these activities. These key elements of NASA's 

safety system are identified and delineated as follows: 

Industrial Safety. This element includes those 

activities which, on a continuing basis, provide protection for the well being of personnel and 

prevention of damage to property involved in I!ASA's 

business and exposed to potential hazards 

associated with carrying out this business. 

Industrial safety relates especially to the 

operation of facilities in the many programs of 

research, developnent, manufacture, test, opera- tion, and niaintenance. Industrial safety 

activities include, but are not limited to, such 

functions as: 

(a) Determination of industrial safety criteria. 

(b) Establishment and implementation of safety 

standards and procedures for operation and 

maintenance of facilities, especially test 

and hazardous environment facilities. 

(c) Development of safety requirements for the 

(d) Establishment and implementation of safety 

design of new facilities. 

standards and procedures for operation of 

program support and administrative aircraft. 

(2) Systems Safety. This element includes those 

activities specifically organized to deal with the 

potential hazards of complex F.&D systems that 

involve many highly specialized areas of technology. It plac'es particular emphasis on achieving safe operation of these systems over their life cycles, and it covers major systems 

for aeronautical and space flight activities, 

manned or unmanned, including associated groundbased research, development, manufacturing, and 

test activities. Systems safety activities 

include, but are not limited to, such functions as : 

(a) Determination of systems safety criteria, 

including criteria for crew safety. 

(b) Determination of safety data requirements. 

(c) Perfornance of systems safety analyses. 

1-11 
Eecember 7, 1967 

103 

NMI 1156.14 

(d) Establishment and implementation of systems safety plans. 

(3) ?ublic Safety. This element includes those 

activities which, on a continuing basis, provide 

protection for the xell being of people and 

prevention of darnage to property not' involved in 

UASA's business, but which may nevertheless be 

exposed to potential hazards associated with carryin% out this business. Public safety activities 

include, but are not limited to, such functions as: 

(a) Determination of public safety criteria. 

(b) Establishment and control of public safety 

hazards associated with facility and systems 

tests and operations. 

(c) Establishment and imnlementation, as required, 

of emergency or catastrophe control plans. 

(4) Safety I4anagement. This element includes both the 

program and functional organizations of I'iASA and 

its contractors involved in the identification of 

potential hazards and their elimination or control as set forth in the foregoing description of 

safety activities. It also includes the uanagement 

systems for planning, inpl.ementing, coordinating, 

and controlling these activities, These mannEe1:ient 

systems include, but me not liniited to, the 

following : 

(a) The authorities, responsibilities, and working 

relationships of the organizations involved 

in safety activities, and the assessment of 

their effectiveness. 

(b) The procedures for insuring the currency and 

continuity of safety activities, especially 

systems safety activities which may extend 

over long periods of time and where management responsibilities are transferred during 

the life cycles of the systems. 

investigations, including those for the followup on corrective actions and the feedback of 

accident/incident information to other 

involved or interested organizations. 

(c) The plans and procedures for accident/incident 

(d) The analysis and dissemination of safety data. 

1-12 
NMI 1156.14 

104 

December 7, 1967 

4. PFiOCEDURES 

2. The Panel will function in an ajvisory ca.pacity to the 

Admi ni s t r it t or , ani; , t h rou h hirn, to those orgenizational 

elenients responsible for ana;;ement of the iJASA shfety 

activiLies. 

b, 'i'he Tenel will be proviciec'. :.:ith all inforination require; 

tc &ischarre its aivisory responsibilities as they 

pertain to both ::IXA and its contrectors' safety 

activities. This inforr)!ation xi11 be made available 

throusii the i:iechanis>:; of a;progriate recorts, End by 

nezns of in situ revie:.is of silr*ety activities at the 

vcricus :i.Gfi anti contractcr sites , i.s dee!,lec; necessary 

by the ?anel zni arranyed t;-.rou;h the Adrministrztor. 

The ?anel \;ill thus be enzbled to cxanine and evaluate 

not cnly the :;enera1 status of the :TASA scfety system, 

but also the key e1e:ner:ts of the Flanned and on-soing 

activities in this system. 

-- 

5. ORGAXIZATIOM 

a. ;Ienbership 

(1) 'The Panel will consist of a maxinun of nine re:nbers, 

who will be appointed by the Adxinistrator. 

Appointi.ients )!ill be for a tern of six years, 

except that, in ortier to provide continuity of 

ner.bership, one-third of the !::embers appointed 

originally to the ?anel will be appointed for a term of two years, one-third for a tern of four 

years, and one-third for a tern of six years. 

(2) Not nore than four r.ienbers of the ?ape1 shall be 

employees of ;.!ASP., nor shall such i:hSA ,:e:.,bers 

constitute a najority of the conposition of the 

Panel at any Liven time. 

members shall be as specified in Section 6 of the 

IJASA Authorization Act, 1968. 

(3) Corlipensation and travel allowances for ?anel 

b. Officers 

(1) The Officers of the Panel shall be a Chairman Lni 

a Vice Chairman, who shall be selected by the ranel 

from their memberaship to serve for one-year terns. 

shall preside at dll rneetinLs of the Panel anii snall 

have the usual powers of a presidin:: officer. 

(2) The Chairman, or Vice Chairman in his Lbsence, 

1-13 
105 

kcember 7, 1967 

c. Coininittees 

NMI 1156.14 

The ?anel is authorizec; to estr-blish special 

committees, 2s :,ecessar'y and as aj;rovec: by the 

Administrator, to carry OUL specified t&sLs ',$ithi:? 

the scope of duties cf the Panel. 

it,ter activities 'xi11 be considere<: 

an inseparable extension or Panel ectivities, anti 

will be in accordance with ail appliczole i,i*oceiiures and rejulations set forth in this 

Instruction. 

The Chairman of each special com.:ittee shall be 2 

member of the Aerospace Safety Advisory Pacel. Tne 

other committee members ,nay or my not be nembers 

of the Panel, as recominended by the Panel anu approved by the Administrator. 

Appointment of Panel i ~ n b e r s t o comicittees as officers or ineinbers will be either for one year, 

for the duration of their term as Fanel r.er.bers, or for the lifetime of the comittee, whichever is the 

shortest. Appointzents of non-Panel members to 

coniinittees will be for a period of one year or for 

the lifetime of' the coniflittee, whichever is shorter. 

Compensation and travel allowances for comrittee 

members who are not members of the Panel'shell be 

the same as for members of the Panel itself, excert 

that conpensation for such coi~iinittee niembers 

appointed from outside the Federal Governiiient shall 

be at the rate prescribed by the Administrator for 

comparable services, 

d. Meetings 

Regular meetings of the Panel will be held as often 

as necessary and at least twice apar. One meeting 

each year shall be an Annual !*leetine. Business 

conducted at this meeting will include selecting 

the Chairmanmd the Vice Chairman of the Panel, 

recommending new comittees and comnittee members 

as required or desired, approving the Panel's 

annual report to the Administrator, and such other 

business as may be required. 

Special meetings of the Panel may be called by the 

Chairinan, by notice served personally upon or by mail or telegraph to the usual address of each 

member at least five days prior to the meeting. 

1-14 
NMI 1156.14 

106 

December 7, 1967 

Special meetinq shall be called in the same manner by the Chairman, upon the written request 

of three menbers of the Panel. 

If practicable, the object of a special meeting 

should be sent in writing to all uenbers, 2nd if 

possible a special meeting should be avoided by 

obtaining the views of members by ma41 or otherwise, 

both on the question requiring the meetina anc; on the question of calling a special iceeting. 

All meetings of special conr;.ittees will be called 

by their respective c.hairsien pursuant to and in 

accordance with performing their specifieii tasks. 

Flinutes of all meetings of the Panel, and of s;.ecial 

committees established by the Panel, ,,,!ill be kepr;. 

Such minutes shall, at a mininun, contain a recorf. 

of persons present, a description of lnatters discussed and conclusions reached, and copies of all 

reports received, issued, or approved by the Panel or committee. The accuracy of all minutes will be 

certified to by the Chairman of the Panel (or by 

the Vice Chairman in his absence) or of the 

comni t t ee . e. Reports and Records 

(1) The Panel shall submit an annual report to the 

(2) The Panel will submit to the Administrator reports 

Administrator. 

on all safety reviews and evaluations with coments 

and recommendations as deemed appropriate by the 

Panel, 

(3) All records and files of the Panel, including 

agendas, minutes of Panel and cormittee meetings,' 

studies, analyses, reports, or other data coinpilations or work papers, made available to or prepared by or for the Panel, will be retained by 

the Panel. 

f. Avoidance of Conflicts of Interest 

(1) Nongovernmental menbers of the Panel, and of 

special committees establishea by the Panel, are "Special Government Zm?ioyees" within the cieaning 

of NHB 1900.2A, which sets fo?th guiciance to LASA 

Special Government 5rr.pioyees rezarding the 

avoidance of conflicts of interest and the 

observance of ethicel stantiards of conduct. A 

1- 15 
107 

December 7, 1967 NMI 1156.14 

copy of NHB 1900.2A and related. IikSA instructions 

on conflicts of interest will be furnished to each 

Panel or committee meraber at the time of his 

appointment as a NASA consultant or expert. 

committee will submit a "NASA Special Government 

Employees Confidential Sta.tement or" Employment 

and Financial Interests" (XASA Form 1271) prior to 

participatins in the activities of the Panel or 2 

special committee. 

(2) Nongovern::iental members of the Panel or 2 special 

6. SUPPORT 

a. A staff, to be comprised of full-tine NASA employees, 

shall be established to sunnort the Panel. The inembers 

cf this staff will be fully'responsive to direction from 

the Chalrman 01 the Panel. 

b. The director of this staff will serve as Executive 

Secretary to the ?anel. The Zxecutive Secretary of the 

Panel, in accoroance vith tne specific instructions from 

the Chairman of the Panel, shall: 

(1) Adr;inister the affairs of the Panel and have general 

supervision of all arranxements for safety reviews 

and evaluetions, and other natters undertaken by 

the Tanel. 

(2) Insure that a written record is kept of 211 

transactions, and submit the sane to the Panel for 

eggroval at each subsequent meeting. 

special committees of the Panel. 

(3) Insure that the sage service is provided for 311 

CFR Title 14, Chapter 5, Subpart 1209.5. 

1-16 
108 

CHAPTER 2 

BOARD HISTORY AND PROCEDURES 
109 

PART 1. SUMMARY OF BOARD RISTORY AND PROCEDURES 

The Ap0110 13 Review Board was established on April 17, 1970, by 

the NASA Administrator and Deputy Administrator under the authority of 

NASA Management Instruction 8621.1, dated April 14, 1966. In the letter 

establishing the Board, Mr. Edgar M. Cortright, Director of Langley 

Research Center, was appointed as Chairman and the general responsibilities of the Board were set forth. The seven additional members of the 

Board were named in a letter from the Administrator and the Deputy 

Administrator to the Chairman, dated April 21, 1970. This letter also 

designated a Manned Space Flight Technical Support official, a Counsel 

to the Board, several other supporting officials, and several observers 

from various organizations. In addition, in a letter dated April 20, 

1970, to Dr. Charles D. Harrington, Chairman of the NASA Aerospace 

Safety Advisory Panel, that Panel was requested to review the Board's 

procedures and findings. 

The Review Board convened at the Manned Spacecraft Center, Houston, 

Texas, on Tuesday, April 21, 1970. Four Panels of the Board were formed, 

each under the overview of a member of the Board. Each of the Panels 

was chaired by a senior official experienced in the area of review 

assigned to the Panel. In addition, each Panel was mamed by a number 

of specialists, thereby providing a nucleus of expertise for the review 

activity. During the period of the Board's review activities, <he 

Chairmen of the four Panels were responsible for the conduct of evaluations, analyses, and other studies bearing on their Panel assignments, 

for preparing preliminary findings and recommendations, and for developing 

other information for the Board's consideration. To overview these 

Panel efforts, each member of the Board assumed specific responsibilities 

related to the overall review. 

In addition to the direct participants in the Board activity, a 

number of observers and consultants also attended various meetings of 

the Board or its constituent Panels. These individuals assisted the 

Review Board participants with advice and counsel in their areas of 

expertise and responsibilities. 

While the Board's intensive review activities were underway, the 

Manned Spacecraft Center Apollo 13 Investigation Team, under James A. 

McDivitt, Colonel, WAF, was also conducting its own analysis of the 

accident on Ap0110 13. Coordination between the Investigation Team 

work and the Ap0110 13 Review Board activities was effected through the 

MSF Technical Support official and by maintaining a close and continuing working relationship between the Panel Chairmen and officials of 

the MSC Investigation Team. 

2-1 
110 

The Board Chainnm established a series of administrative procedures 

to guide the Board's activities. In addition, specific assignments of 

responsibility were made to all individuals involved in the Board's 

activities so as to insure an efficient review activity. Overall logistic and administrative support was provided by MSC. 

The Board conducted both Executive and General Sessions. During 

the Executive Sessions, plans were agreed upon for guiding the Board's 

activities and for establishing priorities for tests, analyses, studies, 

and other Board efforts. At the General Sessions, status of Panel 

activities was reviewed by the Board with a view towards coordination 

and integration of all review activities. In addition, Board members 

regularly attended daily status meetings of the Manned Spacecraft Center 

Investigation Team. 

In general, the Board relied on Manned Spacecraft Center postmission 

evaluation activities to provide the factual data upon which evaluation, 

assessment, and analysis efforts could be based. However, the Board, 

through a regular procedure, also levied specific data collection, reduction, and analysis requirements on MSC. Test support for the Board 

was conducted primarily at MSC but also included tests run at other 

NASA Centers. Members of the Board and its Panels also visited a number 

of contractor facilities to review manufacturing, assembly, and test 

procedures applicable to the Apollo 13 mission. 

The Chairman of the Board provided the NASA Deputy Administrator 

with oral progress reports. These reports summarized the status of 

Review Board activities at the time and outlined the tasks still ahead. 

All material used in these interim briefings was incorporated into the 

Board's official files. 

As a means of formally transmitting its findings, determinations, 

and recommendations, the Board chose the format of this Final Report 

which includes both the Board's judgments as well BS the reports of the 

individual Panels. 

A general file of all the data and information collected and examined 

by the Board has been established at the Langley Research Center, Hampton, 

Virginia. 

data at MSC. 

In addition, the MSC Investigation Team established a file of 

2-2 
111 

PART 2. BIOGRAPHIES OF BOARD MEMBERS, OBSERVERS, AND PANEL CHAIF" 

CHAIRMAN OF THE APOLLO 13 REVIEW BOARD 

EXAR M. CORTRIGHT 

NASA Langley Research Center 

Edgar M. Cortright, 46, Director of the NASA Langley Research Center, 

Hampton, Virginia, is Chairman of the Apollo 13 Review Board. 

Mr. Cortright has been an aerospace scientist and administrator for 

22 years, He began his career at NASA's Lewis Research Center, Cleveland, 

Ohio, in 1948 and for the next 10 years specialized in research on highspeed aerodynamics there. 

In October 1958, Mr. Cortright was named Chief of Advanced Technology 

Programs at NASA Headquarters, Washington, D. C., where he directed initial formulation of NASA's Meteorological Satellite Program. 

became Assistant Director for Lunar and Planetary Programs and directed 

the planning and implementation of such projects as Mariner, Ranger, and 

Surveyor, 

In 1960, he 

Mr. Cortright became Deputy Director of the Office of Space Sciences 

in 1961, and Deputy Associate Administrator for Space Science and Applications in 1963, in which capacities he served as General Manager of 

NASA's space flight program using automated spacecraft. He joined the 

Office of Manned Space Flight as Deputy Associate Administrator in 1967 

and served in a similar capacity until he was appointed Director of the 

Langley Research Center in 1968. 

He is a Fellow af the American Institute of Aeronautics and Astronautics and of the American Astronautical Society. He has received the 

.Arthur S. Fleming Award, the NASA Medal for Outstanding Leadership, and 

the NASA Medal for Distinguished Service. 

Mr. Cortright is the author of numerous technical reports and 

articles, and compiled and edited the book, "Exploring Space With a 

Camera. I' 

He is a native of Hastings, Pennsylvania, and served as a U.S. Navy 

officer in World War 11. He received Bachelor and Master of Science 

degrees in aeronautical engineering from the Rensselaer Polytechnic 

Institute. 

Mr. and Mrs. Cortright are the parents of two children. 

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112 

MEMBERS OF THE PPOLLO 13 REVIEW BOARD 

ROBERT F, ALLNUTI 

NASA Headquarters 

Robert F. Allnutt, 34, Assistant to the NASA Administrator, 

Washington, D. C., is a member of the Apollo 13 Review Board. 

Mr. Allnutt was named to his present position this year. Prior to 

that, he had been Assistant Administrator for Legislative Affairs since 

1967. 

He joined NASA in 1960 as a patent attorney at the Langley Research 

Center, Hampton, Virginia. 

quarters, Washington, D. C. 

In 1961, he was transferred to NASA HeadMr. Allnutt served as Patent Counsel for Communications Satellite 

Corporation from January to September 1965, when he returned to NASA 

Headquarters as Assistant General Counsel for Patent Matters. 

He is admitted to the ?ractice of law in the District of Columbia 

and the state of Virginia and is a member of the American Bar Association 

and the Federal Bar Association. 

Mr. Allnutt was graduated from Virginia Polytechnic Institute with 

a B.S. degree in industrial engineering. He received Juris Doctor and 

Master of Laws degrees from George Washington University Law School. 

Mr. and Mrs. Allnutt are the parents of two sons. The family lives 

in Washington, D. C. 

NEIL A. ARMSTRONG 

NASA Astronaut 

Neil A. Armstrong, 39, NASA astronaut, is a member of the Apollo 13 

Review Board. 

Commander of the Apollo 11 mission and the first man on the Moon, 

Mr. Armstrong has distinguished himself as an astronaut and as an engineering test pilot. 

Prior to joining the astronaut team at the Manned Spacecraft Center, 

Houston, Texas, in 1962, Mr. Armstrong was an X-15 rocket aircraft 

project pilot at the NASA Flight Research Center, Edwards, California. 

2-4 
113 

Mr. Armstrong joined NASA at the Lewis Research Center, Cleveland, 

Ohio, in 1955, and later transferred to the Flight Research Center as an 

aeronautical research pilot. 

His initial space flight was as command pilot of Gemini VIII, 

launched March 16, 1966. 

two vehicles in space. The flight was terminated early due to a malfunctioning thruster. and the crew was cited for exceptional piloting 

skill in overcoming the problem and accomplishing a safe landing. He 

has served on backup crews for both Gemini and Apollo. 

He performed the first successful docking of 

Mr. Armstrong is a Fellow of the Society of Experimental Test 

Pilots, Associate Fellow of the American Institute of Aeronautics and 

Astronautics, and member of the Soaring Society of America. He has received the Institute of Aerospace Sciences Octave Chanute Award, the 

AM Astronautics Award, the NASA Exceptional Service Medal, the John F. 

Montgomery Award, and the Presidential Medal of Freedom. 

He is a native of Wapakoneta, Ohio, and received a B.S. degree in 

aeronautical engineering from Purdue University and a M.S. degree from 

the University of Southern California. He was a naval aviator from 

1949 to 1952 and flew 78 combat missions during the Korean action. 

Mr. and Mrs. Armstrong have two sons, 

JOHN F. CLARK 

NASA Goddard Space Flight Center 

Dr. John F. Clark, 49, Director of the NASA Goddard Space Flight 

Center, Greenbelt, Maryland, is a member of the Apollo 13 Review Board. 

He is an internationally known authority on atmospheric and space 

sciences, holds four patents in electronic circuits and systems, and has 

written many scientific papers on atmospheric physics, electronics, and 

mathematics. 

Dr. Clark joined NASA in 1958 and served in the Office of Space 

Flight Programs at NASA Headquarters until 1961 when he was named 

Director of Geophysics and Astronomy Programs, Office of Space Sciences, 

From 1962 until 1965, he was Director of Sciences and Chairman of the 

Space Science Steering Committee, Office of Space Science and Applications. 

In 1965, Dr. Clark was appointed Deputy Associate Administrator for 

Space Science and Applications (Sciences), and later that year, Acting 

Director of Goddard. He was named director of the center in 1966. 

2-5 
114 

Dr. Clark began his career in 1942 as an electronics engineer at 

the Naval Research Laboratory, Washington, D.C. From 1947 to 1948 he 

was Assistant Professor of Electronic Engineering at Lehigh University, 

Bethelem, Pennsylvania. He returned to NRL in 1948; and prior to joining NASA, served as head of the Atmospheric Electricity Branch there. 

He is a member of the American Association of Physics Teachers, 

American Geophysical Union, Scientific Research Society of America, 

Philosophical Society of Washington, the International Scientific Radio 

Union, and the Visiting Committee on Physics, Lehigh University. He 

received the NASA Medals for Exceptional Service, Outstanding Leadership, 

and Distinguished Service. 

Dr. Clark was born in Reading, Pennsylvania. He received a B.S. 

degree in electrical engineering from Lehigh University, M.S. degree in 

mathematics from George Washington University, and Ph. D. in physics 

from the University of Maryland. 

Dr. and Mrs. Clark have two children and live in Silver Springs, 

Maryland. 

WALTER R. HEDRICK, JR. 

Headquarters , USAF 

Brig. Gen. Walter R. Hedrick, Jr., h8, Director of Space, Office 

of the Deputy Chief of Staff for Research and Development, Headquarters, 

USAF, Washington, D.C., is a member of the Apollo 13 Review Board. 

He has participated in most of the Air Force's major nuclear test 

projects and has extensive experience as a technical project officer 

and administrator. 

General Hedrick joined the Army Air Corps as an aviation cadet in 

1941 and flew in combat with the 86th Fighter Bomber Group during 

World War 11. After the War, he was assigned to the 19th Air Force, the 

14th Air Force, and as a project officer under Air Force Secretary 

Stuart Symington. From 1952 to 1955, he was assigned to the Air Force 

Office of Atomic hergy. 

In 1955, he was assigned to the Technical Operations Division, Air 

Force Special Weapons Command, Kirtland Air Force Base, New Mexico. In 

1957, he was named Commander of the 4951st Support Squadron, hiwetok; 

and the following year, he was reassigned to Kirtland AFB as Assistant 

to the Group Commander and later as Air Commander of the 4925th Test Group. 

General Hedrick joined the Special Systems Office, Air Force 

Ballistics Division, Los Angeles, in 1960. He was named Commander of 

2-6 
115 

the Satellite Control Facility in 1965, and in 1966, he was appointed 

Deputy Commander, Air Force Systems Command. He received his present 

assignment in 1967. 

General Hedrick is a Command Pilot and has received numerous Air 

Force awards. 

His home town is Fort Worth, Texas, and he attended Texas Technological College, Lubbock, prior to joining the service. 

B.S. and M.S. degrees in physics from the University of Maryland. 

General and Mrs. Hedrick are the parents of two sons. 

He received 

VINCENT L. JOHNSON 

NASA Headquarters 

Vincent L Johnson, 51, Deputy Associate Administrator for Space 

Science and Applications (Engineering), NASA Headquarters, is a member 

of the Apollo I3 Review Board. 

Mr. Johnson was appointed to his present position in 1967. Prior 

to that time, he had been Director of the Launch Vehicle and Propulsion 

Programs Division, Office of space Science and Applications, since 1964. 

He was responsible for the management and development of the light and 

medium launch vehicles used for NASA's unmanned earth orbital and deep 

space programs. His division also directed studies of future unmanned 

launch vehicle and propulsion system requirements. 

Mr. Johnson joined NASA in 1960, coming from the Navy Department 

where he had been an engineer with the Bureau of Weapons. His first 

assignments with NASA were as Program Manager for the Scout, Delta, and 

Centaur launch vehicles. 

He was a naval officer during World War 11, serving with the Bureau 

of Ordnance. Prior to that, he was a physicist with the Naval Ordnance 

Laboratory. 

Mr. Johnson was born in Red Wing, Minnesota, and attended the 

University of Minnesota. 

He and Mrs. Johnson live in Bethesda, Maryland. They are the 

parents of two children. 

2-7 
116 

MILTON KLEIN 

NASA Headquarters 

Milton Klein, 46, Manager, Space Nuclear Propulsion Office, NASA 

Headquarters, is a member of the Apollo I3 Review Board, 

Mr. Klein has been in his present position since 1967. Prior to 

that he had been Deputy Manager since 1960. 

Office is a joint activity of the Atomic Energy Commission (AEC) and the 

National Aeronautics and Space Administration. The office conducts the 

national nuclear rocket program. He is also Director of the Division of 

Space Nuclear Systems of the AEC, responsible for space nuclear electric 

power activities. 

The Space Nuclear Propulsion 

Mi-. Klein became associated with atomic energy work in 1946, when 

he was employed by the Argonne National Laboratory. In 1950, he joined 

the AEC's Chicago Operations Office as staff chemical engineer. Later, 

he was promoted to Assistant Manager for Technical Operations. Generally 

engaged in reactor development work for stationary power plants, he had 

a primary role in the power reactor demonstration program. 

Mi-. Klein was born in St. Louis, Missouri. He served in the U.S. 

Navy during World War 11. 

He has a B.S. degree in chemical engineering from Washington 

University and a Master of Business Administration degree from Harvard 

University. 

Mr. and Mrs. Klein and their three children live in Bethesda, 

Maryland. 

HANS M. MARK 

NASA Ames Research Center 

Dr. Hans M. Mark, 40, Director of the NASA Ames Research Center, 

Moffett Field, California, is a member of the Apollo 13 Review Board. 

Prior to being appointed Director of the Ames Research Center he 

was, from 1964 to 1969, Chairman of the hpartment of Nuclear Engineering 

at the University of California, Berkeley, California. 

An expert in nuclear and atomic physics, he served as Reactor 

Administrator of the University of California's Berkeley Research 

Reactor, professor of nuclear engineering and a research physicist at 

the University's Lawrence Radiation Laboratory, Livermore, California, 

2-8 
117 

and consultant to the U.S. Army and the National Science Foundation. 

He has written many scientific papers. 

Except for 2 years as an Assistant Professor of Physics at the 

Massachusetts Institute of Technology from 1958 to 1960, Dr. Mark's 

administrative, academic, and research career has been centered at the 

University of California (Berkeley). 

Dr. Mark received his A.B. degree in physics from the University 

of California, Berkeley, in 1951, and returned there as a research 

physicist in 1955, one year after receiving his Ph. D. in physics 

from M.I.T. 

He is a Fellow of the American Physical Society and a member of the 

American Geophysical Union, the American Society for Engineering Education and the American Nuclear Society. 

Dr. Mark was born in Mannheim, Germany, and came to the United 

States when he was 11 years old. He became a naturalized U.S. citizen 

in 1945. 

Dr. and Mrs. Mark are the parents of two children. 

COUNSEL TO THE AF'OLLO 13 REVIEW BOARD 

GEORGE T. MALLEY 

NASA Langley Research Center 

George T. Malley, 57, Chief Counsel, Langley Research Center, 

Hampton, Virginia, is the Legal Counsel to the Apollo 13 Review Board. 

He also served as Counsel to the Apollo 204 Review Board. 

Mr. Malley is the Senior Field Counsel of NASA and has been assigned 

to Aangley since 1959. He was with the Office of the General Counsel, 

Department of the Navy, from 1950 to 1959, where he specialized in 

admiralty and international law. 

He is a retired Navy officer and served on active duty from 1939 to 

1946, mainly in the South Pacific. 

officer of the U.S.S. Fentress. 

His last assignment was commanding 

Mr. Malley has an A.B. degree from the University of Rochester and 

an LL.B. degree from Cornell University Law School. He is a native of 

Rochester, New York, and is a member of the New York Bar and the Federal 

Bar Association. 

Mr. and Mrs. Malley and their two children live in Newport News, 

Virginia. 

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118 

MANNED SPACE FLIGHT TECHNICAL SUPPORT 

CHARLES W. MATHEWS 

NASA Headquarters 

Charles W. Mathews, 49, Deputy Associate Administrator for Manned 

Space Flight, NASA Headquarters, Washington, D. C., directs the Office 

of Manned Space Flight technical support to the Apollo 13 Review Board. 

Mr. Mathews has been a research engineer and project manager for 

NASA and its predecessor, the National Advisory Committee for Aeronautics 

(NACA), since 1943. 

manager of manned space flight. 

In his present assignment, he serves as general 

Prior t o h i s appointment to this position in 1968, he had been 

Director, Apollo Applications Program, NASA Headquarters, since 

January 1967. 

Mr. Mathews was Gemini Program Manager at the Manned Spacecraft 

Center, Houston, Texas, from 1963 until 1967. Prior to that time, he 

was Deputy Assistant Director for Engineering and Development and Chief 

of the Spacecraft Technology Division at MSC. 

Mr. Mathews transferred to MSC (then the Space Task Group) when 

Project Mercury became an official national program in 1958. as Chief of the Operation Division. He had been at the Langley Research 

Center, Hampton, Virginia, since 1943 engaged in aircraft flight research 

and automatic control of airplanes. He became involved in manned spacecraft studies prior to the first Sputnik flights, and he conducted early 

studies on reentry. Mr. Mathews was chairman of the group which developed 

detailed specifications for the Mercury spacecraft. 

He served 

Mr. Mathews has been awarded the NASA Distinguished Service Medal 

and the NASA Outstanding Leadership Medal. He has received the NASA 

Group Achievement Award - Gemini Program Team. 

He is a Fellow of the American Astronautical Society and an Associate 

Fellow of the American Institute of Aeronautics and Astronautics. He is 

the author of numerous technical articles published by NASA. 

Mr. Mathews, a native of Duluth, Minnesota, has a B.S. degree in 

aeronautical engineering from Rensselaer Polytechnic Institute, Troy, 

New York. 

Mr. and Mrs. Mathews live in Vienna, Virginia. They have two 

children. 

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119 

APOLLO 13 REVIEW BOARD OBSERVERS 

WILLIAM A. ANDERS 

National Aeronautics and Space Council 

William A. Anders, 36, Executive Secretary, National Aeronautics 

and Space Council, Washington, D.C., is an official observer of the 

Apollo 13 Review Board. 

Prior to being appointed to his present position in 1969, Mr. Anders 

was a NASA astronaut and an Air Force lieutenant colonel. He was lunar 

module pilot on the Apollo 8 lunar orbital mission, man's first visit 

to the vicinity of another celestial body. 

Mr. Anders joined the NASA astronaut team at the Manned Spacecraft 

Center, Houston, Texas, in 1963. In addition to his Apollo 8 flight, he 

served as backup pilot for Gemini 11 and backup command module pilot for 

Apollo 11, the first lunar landing mission. 

Mr. Anders was commissioned a second lieutenant in the Air Force 

upon graduation from the U.S. Naval Academy. After flight training, he 

served as a pilot in all-weather interceptor squadrons of the Air Defense 

Command. Prior to becoming an astronaut, he was a nuclear engineer and 

instructor pilot at the Air Force Weapons Laboratory, Kirtland Air Force 

Base, New Mexico. 

He is a member of the American Nuclear Society and has been awarded 

the Air Force Commendation Medal, Air Force Astronaut Wings, the NASA 

Distinguished Service Medal, and the New York State Medal for Valor. 

Mr. Anders was born in Hong Kong. He received a B.S. degree from 

the U.S. Naval Academy and an M.S. degree in nuclear engineering from 

the Air Force Institute of Technology. 

Mr. and Mrs. hders are the parents of five children. 

CHARLES D. HARRINGTON 

Douglas United Nuclear, Inc; 

Dr Charles D Harrington, 59, President and General Manager, 

Douglas United Nuclear, Inc., Richland, Washington, is an official 

observer of the Apollo 13 Review Board. 

Dr. Harrington, who has been associated with all phases of the 

chemical and nuclear industrial fields since 1941, is Chairman of the 

Aerospace Safety Advisory Panel, a statutory body created by Congress. 

2-11 
From 1941 to 1961, he was employed by the Mallinckrodt Chemical 

Works, St. Louis, Missouri. Dr. Harrington started with the company 

as a research chemist and in 1960, after a procession of research and 

management positions, was appointed Vice President, Mallinckrodt Nuclear 

Corporation and Vice President, Mallinckrodt Chemical Works. 

In 1961, when the fuel material processing plant of Mallinckrodt 

became the Chemicals Division of United Nuclear Corporation, Dr. Harrington 

was named Vice President of that division. 

He became Senior Vice President, United Nuclear Corporation, 

Centreville, Maryland, in 1963, 

In 1965, Dr. Harrington was appointed President and General Manager, 

Douglas United Nuclear, Inc. The company manages production reactors 

and fuels fabrication facilities at Hanford, Washington, for the Atomic 

Energy Commission, 

He is the co-author of a book, "Uranium Prodilction Technology," and 

has written numerous technical papers. He has received the Mid-West 

Award of the American Chemical Society for contributions to technology 

in the nuclear energy field. 

He is director of several corporations, including United Nuclear, 

as well as professional councils and societies. 

Dr. Harrington has M.S., M.A., and Ph. D. degrees in chemistry from 

Harvard University. 

I. IRVING PINKEL 

NASA Lewis Research Center 

I. Irving Pinkel, 57, Director, Aerospace Safety Research and Data 

Institute at the NASA Lewis Research Center, Cleveland, Ohio, is an 

official observer of the Apollo 13 Review Board. 

Until recently, he directed research at Lewis Research Center on 

rocket propellant and electric power generation systems for space 

vehicles, compressors and turbines for advanced aircraft engines, and 

lubrication systems for rotating machines for these systems. 

Mr, Pinkel entered Government scientific service in 1935 as a 

physicist with the U.S. Bureau of Mines, Pittsburgh, Pennsylvania. In 

1940, he joined the staff of the Langley Research Center, Hampton, 

Virginia, as a physicist. When the Lewis Research Center was built in 

1942, he transferred there. 

2-12 
He has been elected to Phi Beta Kappa, Sigma Xi, honorary scientific 

society, and Pi Mu Epsilon, honorary mathematics fraternity. He is an 

Ohio Professional Engineer, served on the former NACA suzcommittees on 

Meteorological Problems, Icing Problems, Aircraft Fire Prevention and 

Flight Safety, and is a member of the NASA Research and Technology Advisory Subcommittee on Aircraft Operating Problems. He has been a Special 

Lecturer, Case Institute of Technology Graduate School. 

Mr. Pinkel has received the Flight Safety Foundation Award for contributions to the safe utilization of aircraft, the Laura Taber Barbour 

Award for development of a system for suppressing aircraft crash fires, 

the NACA Distinguished Service Medal, and the NASA Sustained Superior 

Performance Award. 

He was born in Gloversville, New York, and was graduated from the 

University of Pennsylvania. 

Mr. and Mrs. Pinkel live in Fairview Park, Ohio. They are the 

parents of two sons. 

JAMES E. WILSON, JR. 

Committee on Science end Astronautics 

United States House of Representatives 

James E. Wilson, Jr., 39, Technical Consultant, United States House 

of Representatives Committee on Science and Astronautics, is an official 

observer of the Apollo 13 Review Board. 

Mr. Wilson has been technical consultant to the Committee since 

1963, 

U.S. Naval Propellant Plant, Indian Head, Maryland. Mr. Wilson managed 

the Polaris Program at Indian Head from 1956 to 1961. 

From 1961 to 1963, he was Director of Research and Development, 

From 1954 to 1956, Mr. Wilson served as an officer in the U.S. Army 

Signal Corps. He was a development engineer with E. I. Wont, Wilmington, 

Delaware, from 1953 to 1954. 

Mr. Wilson is a member of Phi Sigma Alpha, a National Honor Society; 

American Institute of Chemical Engineers; American Chemical Society; and 

American Ordnance Association. 

Mr. Wilson is co-author of several publications of the House Committee on Science and Astronautics. 

He received a B.S. degree in chemical engineering from the University of Maine and a Master of Engineering Administration degree from 

George Washington University. 

2-13 
Mr. and Mrs. Wilson live in LaPlata, Maryland. They have two 

children. 

APOLLO 13 REVIEW BOARD PANEL CHAIRMEN 

SEYMOUR C. HIMMEL 

NASA Lewis Research Center 

Dr. Seymour C. Himmel, Assistant Director for Rockets and Vehicles, 

Lewis Research Center, Cleveland, Ohio, heads the Design Panel of the 

ApoYo 13 Review Board. 

Dr. Himmel joined Lewis in 1948 as an aeronautical research scientist. He has occupied supervisory positions since 1953. 

He has been awarded the NASA Exceptional Service Medal and the NASA 

Group Achievement Award as manager of the Agena Project Group. 

has served on a number of advisory committees. He is an Associate Fellow 

of the American Institute of Aeronautics and Astronautics, and a member 

of Tau Beta Pi and Pi Tau Sigma. He is the author of more than 25 technical papers. 

Dr. Himmel 

Dr. Himmel has a Bachelor of Mechanical Engineering degree from the 

College of the City of New York and M.S. and Ph. D. degrees from Case 

Institute of Technology. 

Dr. and Mrs. Himmel live in Lakewood, Ohio. 

EDWlN C. KILGORE 

NASA Langley Research Center 

Edwin C. Kilgore, 47, Deputy Chief, Engineering and Technical Services, Langley Research Center, Hampton, Virginia, heads the Project 

Management Panel of the Apollo lJ Review Board. 

Mr. Kilgore joined the Langley science staff in 1944 and served in 

a variety of technical and management positions until promotion to his 

present position in 1968. 

He has received the Honorary Group Achievement Award for his role 

in achieving a record of 97 consecutive successes for solid propellant 

rocket motors and the NASA-Lunar Orbiter Project Group Achievement Award 

for outstanding performance. He is a member of Pi Tau Sigma, honorary 

mechanical engineering society. 

2-14 
Mr. Kilgore was born in Coeburn, Virginia. 

Virginia Polytechnic Institute with a B.S. degree in mechanical engineering. 

He was graduated from 

Mr. and Mrs. Kilgore and their two daughters live in Hampton. 

HARRIS M. SCHURMEIER 

California Institute of Technology Jet Propulsion Laboratory 

Harris M. Schurmeier, 45, Deputy Assistant Laboratory Director for 

Flight Projects, California Institute of Technology Jet Propulsion Laboratory, Pasadena, California, heads the Manufacturing and Test Panel 

of the Apollo 13 Review Board. 

Mr. Schurmeier was appointed to his current position in 1969. Prior 

to that he was Mariner Mars 1969 Project Manager, Voyager Capsule System 

Manager and Deputy Manager of the Voyager Project, and Ranger Project 

Manager at JPL. 

He has received the NASA Medals for Exceptional Scientific Achievement and Exceptional Service. In addition, he has received the Astronautics Engineer Award, and the NASA Public Service Award. 

He was born in St. Paul, MiMeSOta. He has received a B.S. degree 

in mechanical engineering, M.S. degree in aeronautical engineering, and 

a professional degree in aeronautical engineering from the California 

Institute of Technology. 

Mr. Schurmeier was a naval officer in World War 11. He and his 

wife and four children live in Altadena, California. 

FRANCIS B. SMITH 

NASA Headquarters 

Francis B. Smith, 47, Assistant Administrator for University Affairs, 

NASA Headquarters, is leader of the Mission Events Panel of the Apollo 13 

Review Board. 

Mr. Smith has been in his present position since 1967. Prior to 

that he had been Assistant Director, Langley Research Center, Hampton, 

Virginia, since 1964. He joined the Langley science staff in 1947. He 

is an expert in several fields, including radio telemetry, radar, electronic tracking systems, and missile and range instrumentation. 

2-15 
Mr. Smith was born in Piedmont, South Carolina, and received a B.S. 

degree in electrical engineqring from the University of South Carolina, 

where he was elected to Phi Beta Kappa. He remained at the University 

as an instructor from 1943 to 1944 and then served in the U.S. Navy until 

1946. 

Mr. and Mrs. Smith and their three children live in Reston, Virginia. 

2-16 
PART 3. BOARD ORGANIZATION AND GENERAL ASSIGNMENTS FOR BOARD PANELS 

BOARD ORGANIZATION 

After reviewing the scope of the Board's charter, the Chairman and 

Board Members agreed upon the Panel and Support Office structure depicted 

on the following organization chart. Each Panel was assigned specific 

responsibilities for reviewing major elements of the overall Board task, 

with particular emphasis upon establishing a sound and independent 

technical data base upon which findings, determinations, and recommendations by the Board could be based. The Panels were staffed with individual NASA specialists and established working arrangements with the 

Manned Space Flight line organization personnel working in analogous 

areas, 

The Board's support offices were structured to provide necessary 

staff. logistics, and administrative support without duplication of 

available MSC assistance. 

In addition to this structure, the Board and Panels also utilized 

the special assistance of expert consultants. 

Panel assignments, complete Panel membership, and the official Board 

organization approved by the Chairman are included in this part of the 

Board report. 

2-17 

47-591 0 - 70 - 9 
APOLLO 

REVIEW B

I E. M. CORTRIGHT. CHAIRMA

R. F. ALLNUTI V. 

OBSERVERS 

W. A. ANDERS 

il 1 ;.:;::STRONG M

n

c. D HARRlNGTON -_--_________-__ J. E WILSON. JR W. R. HEDRICK 

I

I 

OMSF 

TECHNICAL SUPPORT 

C W MATHEWS G I MALLEY E P SWIEDA 

I I MANUFACTURING 

h TEST I MISSION EVENIS 

I F B SMITH 

J J WILLIAMS I 

PO51 INCIDENT 

H M SCHURMEIER 

FABRICATION h 

ACCEPTANCE TESTING 

SYSTEM TESllNG 

OUALITY ASSURANCE 
GENERAL ASSIGNMENTS FOR BOARD PANELS 

[AS DOCUMENTED IN THE BOARD'S ADMINISTRATIVE PROCEDUiiES) 

Panel 1 - Mission Events Panel 

It shall be the task of the Mission Events Panel to provide a detailed and accurate chronology of all pertinent events and actions 

leading to, during, and subsequent to the Apollo 13 incident. This 

information, in narrative and graphical time history form, will provide 

the Ap0110 13 Review Boardan official events record on which their 

analysis and conclusions may be based. This record will be published 

in a form suitable for inclusion in the Review Board's official report. 

The Panel will report all significant events derived from telemetry 

records, air-to-ground communications transcripts, crew and control 

center observations, and appropriate documents such as the flight plan, 

mission technique description, Apollo Operation Handbook, and crew checklists. Correlation between various events and other observations related 

to the failure will be noted. Where telemetry data are referenced, the 

Panel will comment as appropriate on its significance, reliability, 

accuracy, and on spacecraft conditions which might have generated the 

data. 

The chronology will consist of three major sections! Preincident 

Events, Incident Events, and Postincident Events. The decision-making 

process leading to the safe recovery, referencing the relevant contingency plans and available alternates, Will be included. 

Preincident Events. - This section will chronicle the progress of 

the flight from the countdown to the time of the incident. All action 

and data relevant to the subsequent incident will be included. 

Incident Events. - This section will cover that period of time beginning at 55 hours and 52 minutes after lift-off and continuing so long 

as abnormal system behavior is relevant to the failure. 

Postincident Events. - This section will document the events and 

activities subsequent to the incident and continuing to mission termination (Splash). Emphasis will be placed on the rationale used on mission 

completion strategf. 

Panel 1 Membership 

Mr. F. B. Smith, Panel Chairman 

Assistant Administrator for University Affairs 

NASA Headquarters 

Washington, D. C. 

2-19 
128 

Dr. Tom B. Ballard 

Aerospace Technologist 

Flight Instrument Division 

Langley Research Center 

Hampton, Virginia 

Mr. M. P. Frank 

Flight Director 

Flight Control Division 

Manned Spacecraft Center 

Houston, Texas 

Mr. John J. Williams 

Director, Spacecraft Operations 

Kennedy Space Center 

Florida 

Mr. Neil Armstrong, Board Member and Panel Monitor 

Astronaut 

Manned Spacecraft Center 

Houston, Texas 

Panel 2 - Manufacturing and Test Panel 

The Manufacturing and Test Panel shall review the manufacturing and 

testing, including the associated reliability and quality assurance 

activities, of the flight hardware components involved in the flight 

failure az determined from the review of the flight data and the analysis 

of the design. The purpose of this review is to ascertain the adequacy 

of the manufacturing procedures, including any modifications, and the preflight test and checkout program, and any possible correlation of these 

activities with the inflight events. 

The Panel shall consist of three activities: 

Fabrication and Acceptance Testing.- This will consist of reviewing 

the fabrication, assembly, and acceptance testing steps actually used 

during the manufacturing of the specific flight hardware elements involved. Fabrication, assembly, and acceptance testing procedures and 

records will be reviewed, as well as observation of actual operations 

when appropriate. 

Subsystem and System Testing.- This will consist of reviewing all 

the flight qualification testing from the completion of the componentlevel acceptance testing up through the countdown to lift-off for the 

specific hardware involved. Test procedures and results will be reviewed 

2-20 
as well as observing specific tests where appropriate. Results of tests 

on other serial number units will also be reviewed when appropriate. 

Reliability and Quality Assurance.- This will be an overview of both 

the manufacturing and testing, covering such things as parts and material 

qualification and control, assembly and testing procedures, and inspection 

and problem/failure reporting and closeout. 

Panel 2 Membership 

Mr. Harris M. Schurmeier, Panel Chairman 

Deputy Assistant Laboratory Director for Flight Projects 

Jet Propulsion Laboratory 

Pasadena, California 

Mr. Edward F. Baehr 

Assistant Chief, Launch Vehicles Division 

Deputy Manager, Titan Project 

Lewis Research Center 

Cleveland, Ohio 

Mr. Karl L. Heimburg 

Director, Astronautics Laboratory 

Marshall Space Flight Center 

Huntsville, Alabama 

Mr. Brooks T. Morris 

Manager, Quality Assurance and Reliability Office 

Jet Propulsion Laboratory 

Pasadena, California 

Dr. John F. Clark, Board Member and Panel Monitor 

Director 

Goddard Space Flight Center 

Greenbelt, Maryland 

Panel 3 - Design Panel 

The Design Panel shall examine the desim of the oxygen and associated systems to the extent necessary to support the theory of failure. 

After such review the Panel shall indicate a course of corrective action 

which shall include requirements for further investigations and/or redesign. In addition, the Panel shall establish requirements for review 

of other Apollo spacecraft systems of similar design, 

2-21 
The Fanel shall consist of four subdivisions: 

Design Evaluation.- This activity shall review the requirements and 

specifications governing the design of the systems, subsystems and components, their derivation, changes thereto and the reasons therefor; and 

the design of the system in response to the requirements, including such 

elements as design approach, material selection, stress analysis, development and qualification test programs, and results. This activizy 

shall also review and evaluate proposed design modifications, including 

changes in operating procedures required by such modifications. 

Failure !<odes and Mechanisms.- This activity shall review the design 

of the systems to ascertain the possible sources of failure m.d the mariner 

in which failures may occur. In this process, they shall attempt to 

correlzte such modes with the evidence from flight and ground test data. 

This shall include considerations such as: energy sources, materials 

compatibility, nature of pressure vessel failure, effects of environment 

and service, the service history of any suspect systems and components, 

and any degradation t'nat may have occurred. 

Electrical.- This activity shall review the design of all electrical 

components associated with the theory of failure to ascertain their 

adequacy. This activity shall also review and evaluate proposea design 

modifications , including changes in operating procedures required by such 

modifications. 

Related Systems.- This activity shall review the design of all 

systems similar to that involved in the Apollo 13 incident with the view 

cc establishing any commonality of design that may indicate a need for 

redesign, They shall also consider the possibility of design modifications to permit damage containment in the event of a failure. 

Panel 3 Membership 

Dr. Seymour C. Himmel, Fanel Chairman 

Assistant Director for Rockets and Vehicles 

Lewis Research Center 

Cleveland, Ohio 

Mi-. William F. Brown, Jr. 

Chief, Strength of Materials Branch 

Materials and Structures Division 

Administration Directorate 

Lewis Research Center 

Cleveland, Ohio 

2-22 
131 

Mr. R. N. Lindley 

Special Assistant to the Associate Administrator for Manned Space Flight 

NASA Headquarters 

Washington, D. C. 

Dr. William R. Lucas 

Director, Program Development 

Marshall Space Flight Center 

Hmtsviile, Alabama 

Mr. J. F. Samders, Jr. 

Project Officer for Command and Service Module 

Office of Manned Space Flight 

NASA Headquarters 

Washington, D. C. 

Mr. Robert C. Wells 

Head, Electric Flight Systems Section 

Vehicles Sranch 

Flight Vehicles and Systems Division 

Office of Engineering and Technical Services 

Langley Research Center 

Hampton, Virginia 

Mr. Vincent L. Johnson, Board Member and Panel Monitor 

Deputy Associate Administrator for Engineering 

Office of Space Science and Applications 

NASA Headquarters 

Washington, D. C. 

Panel 4 - Project Manageaent Panel 

The Project Management Panel will undertake the following tasks: 

1. Review and assess the effectiveness of the management structure employed in Apollo 13 in all areas pertinent to the Apollo 13 

incident. This review will encompass the organization, the responsibilities of organizational elements, and the adequacy of the staffing. 

2. Review and assess the effectiveness of the management systems 

employed on Apollo 13 in all areas pertinent to the Ap0110 13 incident. 

This task will include the management systems employed to control the 

appropriate design, manufacturing, and test operations; the processes 

used to assure adequate communications between organizational elements ; 

the processes used to control hardware and functional interfaces; the 

safety processes involved; and protective security. 

2-23 
3. Review the project management lessons learned from the Apollo 

13 mission from the standpoint of their applicability to subsequent 

Apollo missions. 

Tasks 1 and 2, above, should encompass both the general review of 

the processes used in Apollo 13 and specific applicability to the possible cause or causes of the mission incident as identified by the Board. 

Panel 4 Membership 

E. C. Kilgore, Panel Chairman 

Deputy Chief, Office of Engineering and Technical Services 

Langley Research Center 

Hampton, Virginia 

R. D. Ginter 

Director of Special Programs Office 

Office of Advanced Research and Technology 

NASA Headquarters 

Washington, D. C. 

Merrill H. Mead 

Chief of Programs and Resources Office 

Ames Research Center 

Moffett Field, California 

James B. Whitten 

Assistant Chief, Aeronautical and Space Mechanics Division 

Langley Research Center 

Hampton, Virginia 

Milton Klein, Board Member and Panel Monitor 

Manager, AEC-NASA Space Nuclear Propulsion Office 

Washington, D.C. 

Board Observers 

William A. Anders 

Executive Secretary 

National Aeronautics and Space Council 

Washington, D.C. 

Dr. Charles D. Harrington 

Chairman 

NASA Aerospace Safety Advisory Panel 

Washington, D.C. 

2-24 
133 

I. Irving Pinkel 

Direct or 

Aerospace Safety Research and Data Institute 

Lewis Research Center 

Cleveland, Ohio 

Mr. James E. Wilson 

Technical Consultant to the Committee on Science 

United States House of Representatives 

Washington, D.C. 

d Astror: 

Apollo 13 Review Board Support Staff 

Brian M. Duff 

Public Affairs Officer 

Manned Spacecraft Center 

Houston, Texas 

Gerald J. Mossinghoff 

Director of Congressional Liaison 

NASA Headquarters 

Washington, D.C. 

Edward F. Parry 

Counsel to Office of Manned Space Flight 

NASA Headquarters 

Washington, D. C. 

Raymond G. Romatowski 

Deputy Assistant Director for Administration 

Langley Research Center 

Hampton, Virginia 

Ernest P. Swieda 

Deputy Chief, Skylab Program Control Office 

Kennedy Space Center, Florida 

Consultants to the Board 

Dr. Wayne D. Erickson, Head Dr. Robert Van Dolah 

Aerothermochemistry Branch 

Langley Research Center 

Hampton, Virginia 

utics 

Acting Research Director 

Safety Research Center 

Bureau of Mines 

Pittsburgh, Pennsylvania 

2-25 
MSC Support to the Board 

These persons were detailed by MSC to support the Apollo 13 Review 

Board during its review activity at MSC. They are identified by WC 

position title. 

Roy C. Aldridge 

Assistant to the Director of Administration 

Mary Chandler 

Secretary 

Rex Cline 

Technical Writer/Editor 

Evon Collins 

Program Analyst 

Leroy Cotton 

Equipment Specialist 

Maureen Cruz 

Travel Clerk 

Janet Harris 

Clerk Stenographer 

Marjorie Harrison 

Secretary 

Phyllis Hayes 

Secretary 

William N. Henderson 

Management Analyst 

Sharon Laws 

Secretary 

Carolyn Lisenbee 

Secretary 

Jamie Moon 

Technical Editor 

Dorothy Newberry 

Administrative Assist ant 

Lettie Reed 

Editorial Assistant 

Charlene Rogozinski 

Secretary 

Joanne Sanchez 

Secretary 

Billie Schmidt 

Fnployee Development Specialist 

Frances Smith 

Secretary 

George Sowers 

Management Presentations Officer 

Elaine Stemerick 

Secretary 

Mary Thompson 

Administrative Assist ant 

Alvin C. Zuehlke 

Electrical Engineer 

Judy Miller 

Secretary 

2-26 
PART 4. SWY OF BOARD ACTrVITlES 

APRIL 19, 197C 

Chairman E. M. Cortright met with Langley officials to begin planning 

the Apollo 13 Review Board approach. 

other specialists were developed for consideration. 

Tentative list of Panel Members and 

APRIL 20, 1970 

Chairman Cortright met with the NASA Administrator, Deputy Administrator, and key NASA officfals in Washington, D.C., to discuss Board 

membership. 

The Chairman met with NASA Office of Manned space Flight top offictals while enroute to MSC on NASA aircraft and discussed program organization plans for review of the accident, and coordination with Apollo 13 

Review Board activity. 

APRIL 21, 1970 

Chairman Cortright met with MSC officials to discuss Apollo 13 

Review Board support. 

A formal MSC debriefing of the Apollo 13 crew was conducted for MSC 

officials and Apollo 13 Review Board personnel already at MSC. 

Detailed discussions between early arrivals on the Review Board and 

the MSC Investigation Team were held to provide quick-look data on the 

Apollo 13 accident and to develop detailed procedures for MSC suaport of 

the Apollo 13 Board. 

Chairman Cortright met with members of the Press to repcrt on early 

activity of the Board and to inform them of plans for keeping the Press 

current on Board activities. 

The first meeting of the Board was held at 8 p.m. to discuss Board 

composition, structure, assignments, and scope of review. Preliminary 

plans were developed for appointing various specialists to assist the 

Board in its analysis and evaluation. 

2-27 
136 

APRIL 22, 1970 

?he Board met with Colonel McDivitt's MSC Investigation Team to review the progress made by :4SC in identifying causes of the accident and 

in developing an understanding of sequences and relationships between 

known inflight events. In addition, MSC officials briefed the Board on 

MSC Investigation Team structxe and assignments. 

The Board met with Panel 1 of the MSC Investigation Team for detailed discussion of inflight events and consideration of early conclusions on implications of preliminary data analysis. 

The Board held its second meeting to disc.Jss MSC investigative 

efforts and additional sppointments of Panel specialists. 

Board members attended Panel 1 evening roundup of day's evaluation 

activities, wnich included detailed discussions of specific studies, 

data rechcticns, and support test activities already underway. 

APRIL 23, 1970 

The Apoilo 13 Review Board established itself in proximity to the 

MSC Investlgation Team in Building 45, and arranged for all admtnistrative and logistics support to the Board. 

A daily schedule of meetings, reviews, briefings, and discussions 

was established, including preliminary plans for contractor meetings, 

special support tests, and accumulation of accident-related informaticn. 

Initial task assignments and responsibilities were made to Board 

Panels as guidance for detailed review work. Individual Board members 

were assigned Panel overview responsibilities or otner special tasks. 

Administrative procedures were developed for Board activity, particularly to provide efficient interface with MSC personnel. 

Board and Panel Members again met with MSC officials to further review the sequence of events in the Apollo 13 mission and to examine early 

hypotheses concerning cames of these events. 

The Board convened for an evening meeting to discuss the progress to 

date and to coordinate Panel activities for the next few days. Discussion 

centered upon immediate requirements for data collection and analysis. 

Chairman Cortright appointed additional NASA specialists in prder to 

bring Panels up to strength. 

2-28 
APRIL 24, 1970 

Board Members, Pane; Chairmen, and MSC officials reviewed additional 

&ata analysis made by MSC and contractor personnel with particular emphasis upon the service module (SM) cryogenic system. 

ke Board convened and reviewed the progress to date. Tentative 

approvals were given for Board trips to North American Rockwell (NR), 

Downey, California, Beech Aircraft, Boulder, Colorado, and other locations. 

Chairman Cortright briefed the Press on progress to date. 

Panel Chairmen and Members continued their detailed analysis of 

failure modes, test histories, mission events, and other data bearing 

upon the accident, 

Board Members and Panel Chairmen met with Mr. Norman mker of NR on 

NR's activities involving design, qualification, and tests of SM cryogenic oxygen tanks. 

APRIL 25, 1970 

'ke Board met to discuss details of onsite inspections of command 

service module (CSM) flight hardware at principal contractor installations. 

Panels examined in detail probable failure modes based on data 

analyzed at that time. 

Specific plans were discussed by the Board relating to evaluation 

of oxygen tank assembly and checkout operations, including review of 

component histories. 

The MSC Investigation Team members briefed Board personnel on 

Kennedy space Center checkout operations of the service module cryogenic 

and electric power systems, including a detailed briefing covering oxygen 

tank detanking operations. 

APRIL 26, 1970 

Board and Panel Members traveled to North American Rockwell, Downey, 

for detailed briefings by NR engineers and management. NR reviewed its 

2-29 
progress in an intensive analysis of the Apollo 13 malfunction, including 

a review of approved special tests. Oxygen tank, fuel cell components, 

assemblies, and ether hardware were also inspected. 

APRIL 27, 1970 

An Executive Session of the Board met to discuss progress of specific 

analyses required to verify tentative conclusions on oxygen tank failure 

and service mcdule EFS failure. 

Additicnal Board specialists arrived at MSC and received detailed 

briefings by MSC and Board personnel on selected aspects of the Apcllo 13 

data. 

Panel Members receiaed and assessed a preliminary MSC evaluation of 

the Apollo 13 accident, including tentative conclusions on the most 

probable failure modes. 

Procedures were established to provide information flow on the status 

of review to Board observers. 

The Board reviewed work plans for the coming week with each Panel and 

established review priorities and special task assignments. 

APRIL 28, 1970 

Chairman Cortright outlined a plan for the Board's preliminary report 

scheduled for presentation to the Deputy Administrator during his visit to 

MSC on May 1. Each Panel Chairman was to summarize the status of his 

Panel's activities for Dr. George Low on Friday, April 29, 1970. 

Board Member Neil Armstrong completed arrangements to provide each 

Board Member and Panel Chairman an opportunity for detailed simulation of 

the Apollo 13 inflight accident using MSC's CSM simulation equipment. 

Board and Panel Members reviewed enhanced photographs of the 

Apollo 13 service module at the MSC Photographic 'Laboratory. 

Dr. von Elbe of Atlantic Research Company briefed Board and Panel 

Members on cryogenics and combustion phenomena. 

A representative of the Manufacturing and Test Panel performed an 

onsite inspection at Beech Aircraft, Boulder. 

2-30 
Manufacture and Test Panel personnel reviewed detanking procedures 

followed at KSC during the Apollo 13 countdown demonstration test (CDDT). 

Board and Panel personnel reviewed progress to date at a general 

Board meeting involving all Review Board personnel. 

APRIL 29, 1970 

Dr. Charles Harrington, Board Observer and Chairman of the Aerospace 

Safety Advisory Panel, arrived for a 2-day detailed review of Board procedures and progress in the accident review. 

The Board reviewed North American Rockwell preliminary recommendatlons Involving oxygen tank redesign. 

The Board continued to review and examine oxygen tank ignition 

so~wces and combustion propagation processes with specialists from MSC, 

other NASA Centers, and contractor personnel. 

The Mission E-rents Panel continued to examine and record details of 

all significant mission events as a basis for other Panel evaluations and 

study. 

Chairman Ccrtright convened two Board meetings to review Panel progress to date and to dlscuss work plans for the next several days. 

The Project !s:anagement Panel visited North American Rockwell at 

Domey to re-riew detailed procedures for acceptance tests, subcontractor 

inspections, project documentation, and other management interface areas. 

APRIL 30, 1970 

Ibe Safety Advisory Panel continued discussions with Board Chairman 

and KSC officials on progress of total Apollo 13 review efforts. 

Panel Members reviewed instrumentation used In Apollo 13 spacecraft 

in order to establish the validity of telemetry data being used in Board 

analysis . 

Chairman Cortright convened two Board meetings to review progress of 

the work and to discuss preliminary findings of the Board. 

Project Management personnel visited Beech Aircraft Corporation to 

review procedures used for assembly of cryogenic oxygen tanks and to discuss communication and information systems within the Apollc Program. 

2-31 
140 

Panels continued to review detailed data in their respective areas. 

MAY 1, 1970 

Board and Panel personnel participated in a joint b!SC/Apollo 13 

Review Board s?atus presentation to the NASA Deputy Adrninistratar. 

meeting co-rered all significant Apollo i3 findings and early conclusions 

sn the cause of the accident and appropriate remediai actions. 

The 

The MSC staff briefed Board Members on i!iitial evaluaticns of proposed design changes in oxygen tank system. 

Panel Members continued to assess data accumulated from the Apollo 13 

mission with particular emphasis upon the design and performance ,-f e;ectric power systems used in the service module. 

Board Members and Panel Chairmen reviewed specific test matrix being 

proposed by Apcllo i3 Review Board specialists covering most significant 

unknowns invol,.-ed in mderstanding failure mect,anisms. 

Board Members met in General Session to discuss preparation of a complete "failure tree" as an additional guide in conducting a complete review and investigation. Specific aspects of this approach were reviewed. 

The Project Management Panel reviewed oxygen tank reliability history 

and quality assurance criteria used In assembly, test, and checkout of 

these systems. 

Panel specialists continued reviewing data from the mission with 

emphasis upon integrating various data points intc logical failure mode 

patterns established by MSC and Board personnel. 

Chairman Cortright and Board Members conducted a detailed review of 

individual Panel status and progress and established milestones for 

additional analytical work and preparation of preliminary findings. 

The Board and Panel agreed to tentative report structure, including 

required exhibits, tables, drawings, and other reference data. 

2-32 
The Board established a system for tabulating all significanL :nission 

events and explanatory data, including the support tests required to 

clarify questions raised by events. 

Panel Members worked on individual analyses with particular attention 

to developing requirements for additional test activity in supporr, of tentative conclusions. 

The Board agreed to strengthen its technical reviews of combustion 

propagation and electrical design by adding specialists in these areas. 

MAY 4, 1970 

The Design Panel continued its intensive review of t'ne "shelf drop" 

incident at NR involving the cryogenic oxygen flight tank used in 

Apollo 13 in order to understand possible results of this even%. 

The Mission Events Panel continued to analyze telemetry data received 

by MSC, with particular attention on data received in proximity to the 

data dropout period during the Apollo 13 mission and on fan turnons during 

the flight. 

The Board transmitted a formal listing of 62 requests for data, 

analyses, and support tests required for Board re-:iew activity. 

The Board continued to meet with individual Panels and support 

offices to review the status of preliminary findings and work completed. 

The Board met in General Session to discuss the scope and conduct of 

support test activity, including careful documentation of test method: and 

application of test results. 

MSC personnel briefed Panel Members on availability of additional 

telemetry data in the MSC data bank in order to insure Board consideration of all possible useful data. 

Panels commenced initial drafting of preliminary findings in specific 

areas, including summary descriptions of system performance during the 

Apollo 13 flight. 

Tne Board met with the MSC Investigation Team for complece review of 

the proposed test program. 

2-33 

47-591 0 - 70 - 10 
142 

Board Members, MSC personnel, and Members of NASA's Aerospace Safety 

Advisory Panel met for detailed discussions and evaluation of accident 

review status and progress. The review covered oxygen tank questions, 

recovery operations. and a mission simulation by MSC astronauts. 

Panel Members continued to work on the preparation of preliminary 

Panel drafts. 

Chairman Cortright transmitted additional requests for tests to MSC 

and modified procedures for control of overall test activity relating to 

the Apollo 13 accident. 

The General Board Session reviewed complete analysis and test support 

activities being conducted for the Board and MSC at various governmental 

and contractor installations. 

Board and Panel Members met to discuss Ames laboratory tests concerning liquid oxygen combustion initiation energies required in the 

cryogenic oxygen tank used in the Apollo 13 SM. 

Panel 1 Members reviewed mission control equipment and operating 

procedures used during the Apollo 13 mission and reviewed actual mission 

events in detail. 

The Panels continued to develop preliminary drafts of their reviews 

and analyses for consideration by the Board. 

MAY 8, 1970 

Dr. Robert Van Dolah, Bureau of Mines, joined the Board as a consultant on combustion propagation and re,viewed Apollo 13 Review Board 

data developed to date. 

The General Board Session convened to review proposed report format 

and scope. An agreement was reached on appendices, on the structure of 

the report, and on the degree of detail to be included in individual Panel 

reports. 

Chairman Cortright assigned additional specific test overview responsibilities to members of the Apollo 13 Review activity. 

2-34 
Panel 1 conducted a formal interview with cke MSC Flight Director 

covering all significant mission events from the standpoint of ground 

controllers. 

Panels 2 through 4 continued developing preliminary reports. Panel 4 

announced a formal schedule of interviews of NSC, contractors, and NASA. 

Headquarters personnel. 

Board Members explored in detail possible failure mode sequences 

developed by MSC personnel involving ignltion and combustion within the 

234 cryogenic oxygen tank. 

The Board recessed for 3 days, leaving a cadre of personnel at MSC 

to edit preliminary drafts developed by the Panels and to schedule fursher 

activity for the week of May 11. 

Board in recess. 

MAY 10, 1970 

Board in recess. 

MAY 11, 1970 

Board in recess. MSC support personnel continued work obtaining 

additional technical data for Board review. 

MAY 12, 1970 

Board Members returned to MSC. 

Board Members attended a General Session to review progress and 

status of the report. 

Panel Chairmen reported on individual progress of work and establlshed schekles for completion of analyses and evaluations. 

Chairman Cortrighs reported on t'ne Langley Researcn Center support 

test program aimed at simulation of SM panel ejection energy pulses. 

2-35 
144 

MAY 13, 1970 

Board Members reviewed preliminary drafts cf report chapter cn Review and Analysis and Panel 1 report on Mission Events. 

Mission Events Panel Members interviewed Electrical, Electronic, and 

Communications Engineer (EECOM) and one of the Apollo 13 Flight Directors 

on activities which took place in the Mission Control Center (MCC) during 

and after the flight accident period. 

Panel 4, Project Eanagernent Panel, conducted interviews with principal Apollo 13 program personnel from MSC and contract organizations. 

Panel Members continued drafting preliminary versions of Panel reports for review by the Board. 

Manufactoring and Test Panel representatives discussed progran tor 

oxygen tank testing to be conducted at Beech Aircraft. 

Board Members met in General Session to re7iew report milestones and 

required test data for the week ahead. 

Board met in General Session to review Panel report progress and to 

agree to firm schedules for completion of all Review Board assignments. 

Project Management Panel continued to interview key Apollo project 

personnel from NASA Centers and contractors. 

Panel Members circulated first drafts of all Panel reports to Board 

Members for review and correction. 

Mission Events Panel personnel interviewed Apollo 13 Command Module 

Pilot John Swigert to verify event chronology compiled by the Panel and 

to review crew responses during Apollo 13 mission. 

Project Management Panel continued interviewing key project personnel 

with NASA Centers and contractors. 

2-36 
:4SC personnel provide Board Members and Panel Chairmen with a detailed briefing on all support tescs and analyses being performed in 

connection with the MSC and Board reviews. 

Board Members met in Executive Session to review preliminary drafts 

cf Panel reports and findings and determinations and to provide additional 

instructions and guidance to Panel Chairmen. 

Panel Members continued to review and edit early Panel drafts and t3 

compile reference data in support of findings. 

Board met in General Session to review further revisions of preliminary findings and determinations and to establish working schedules for 

completion cf the Board report. 

Panel Members continued to edit and refine Panel reports on basis of 

discussions with MSC personnel and further analysis of Apollo 13 documentat ion. 

Draft material for all parts of Board report was reviewed by Panel 

Xembers and staff. Changes were incorporated in all draft material and 

recirculated for additional review and comment. 

Board Members met in General Session to review report progress and 

to examine results from recent support tests and analyses being conducted 

at various Government and contractor instailations. 

The Apollo 13 Iieview Board discussed a continuing series of support 

tests for recommendation to MSC following presentation of report and recess of the Board. 

MAY 18, 1970 

Board Members reviewed Special Tests and Analyses Appendix of the 

report and examined results of completed tests. 

Board met in General Session to discuss control procedures for reproduction and distribution of Board report. 

2-37 
Mission Events Panel distributed a final draft of their report for 

re--ie?r by Bsard Members, 

Board reviewed a preliminary draft of findings and detem.inations 

prepared by Panel Chairmen, Board Members, and Board Chairman. 

X Marxfacture and Test Panel representative re-Tiewed special oxygen 

rank test programs at Eeecn Aircraft. 

3oard :knbers met in 3xecutii.e Sessisn to co-tinue evaluazion an3 

assessxent of preliminary findlngs, determinations, and recommendations 

prepared by individual aoard Members and Panel Chairmen. 

Esa.r3. met in General Session to review final draft of blission E;ienzs 

Panel repzrt. 

I.:anufac",x-e and Test Panel preliminary report was distributed to 

Board ikmbers for re-iiew and comment. 

Design Panel preliminary report was distributed to Board :iIembers for 

re.;iew and corment , 

Design Panel Members met wit'n MSC Team officials to discuss further 

test and analyses nupport for the Boar3. 

Board Members met in Executive Session to review and evaluate reports 

from the Design Panel and from the MarJfacturing and Test Panel. 

Project Management Panel distributed final draft of its report to 

Board Members for review and comment. 

Chairman Cortright met with Mr. Bruce Lundin of the Aerospace Safety 

Ad-risory Panei to discass progress of Eoard re-:iew and analysis. 

MY 21, 1970 

Board Members met in Executive Session for final review of Project 

Nanagement Fanei report, 

2-38 
Board Merhers and others met with NSC officials to review in dezail 

the acti-:ities and actions saken after the Apollo 204 accide:it concerning 

ignition flammability for naferials and control in the CSM. 

A third draft of preliminary findings, determinations, and recommendations was developed and circ;lated bj- the Chaiman for review and 

comment. 

Arrangements were made wish NASA Headquarters officials for packaging, deli-zery, and distribution of the Board's final report. 

Mission Events Panel conducted an interview with Luriar Module Pilot 

:iaise to revie?? selected mission events bearing on the accident, 

MY 22, 1970 

:siissicn E.rents Panel representatives met with MSC officials to review 

in detail se-ieral events which occ-mred during later flight stages. 

8oard met in Executive Session to assess latest drafts of findings, 

determinations, and recommendations circulated by the Chairman. 

Board met in General Session to review total progress in all report 

areas and to establish final schedule for preparation of Board report. 

Larigley Research Center representative M. Ellis briefed the 3oard on 

ignition and combustion of rcaterials in oxygen atmosphere tests being conducted in sup'jort of the Apoilo 13 Iie-liew. 

Board Observer I. I. Pinkel briefed the Board on Lsiris Research 

Center fire propagation tests involving Teflon. 

FL4Y 23, 1970 

3oard Members reviewed Chapter 4 of Board report entitled "Re-riew 

and Analysis. " 

Panel Chairmen reviewed draft findings and determinations prepared 

by the Zoard. 

2-39 
MAY 24, 1970 

Board !<embers reviewed NAS.4 P.erospace Safety 3anei report covering 

Apollc activities during the period of 1968-69. 

Board met in Executive Session for detailed review of support test 

status and progress arid of dccumentation describing the results of test 

activity. 

Board met in Exec.dtive Sesslcn for further review of findings, 

determLnations: and recommendations. 

Board me: in .Zxecui,i-;e Sessicr, to revlew test progress and decided 

to postpone submittal cf final report until June 8 $3 order to consider 

results of Langley Research Center panel ejection tests. 

Board Members continued to review MSC Investigation Team preliminary 

drafys and reflne ApcXo 13 data in the various Board appendices. 

Board met in Txe-ative Session for further consideration of findings, 

determinations, and recommendations. 

MAY 26, 1970 

Board met in Ceneral Session and interviewed Astronaut James Lcvell 

regarding crew understanding of inflight accident. 

Board Members re-iiewed proposed MSC tank combustion test and agreed 

to test methodology and objectives. 

Panel Members continued preparation of individual Tanel reports. 

MAY 27, 1970 

Board and Panel Nembers received a detailed briefing on thermostatic 

switch failure during KSC heater tube temperature tests. 

Aerospace Safety Advisory Panel met with Chairman Cortright, Board 

Members, and Panel Chairmen to review Board progress and status of 

findings and conclusions. 

2-40 
149 

Board met in General Session to review status of Panel reports, 

documentaticn of test data and resiilts, and plans for report typing and 

re-riew. 

Board agreed to recess for several days to acciunulate additional 

zest information on panel separation and full s:ale tank ignition data. 

MAY 28, 1970 

Board in recess. 

K3.Y 29, 1970 

Board in recess. 

3card in recess. 

Board in recess. 

Board Members returned to MSC. 

Board and Panel Members met in General Session to discuss revisions 

of Panel reports in light of latest information regarding thermostatic 

switc:h failure during CDUT at KSC. 

Board approved new schedule for Board report calling for final 

versions of Panel reports by Monday, June 8. 

2-41 
150 

Chairman Cortright briefed the Press on che szatus of tke 3card's 

work and future plans. 

Board and Panel biembers participated in a detailed interview and 

discussion with MSC and contractor personnel regarding specific ccordination steps taken during oxygen tank no. 2 detanking operations at KSC. 

3oard hiembers met in 2xec.acive Session tc re7:iev latesz Gest reslllts 

an5 :o assess szazus of Board findizgs an2 determinations. 

3card an3 Panel Members met with MSC Progrm 3ffice perscmei for a 

Setalled update of recent I6SC information ar.6 analyses stemming fron ongoing test programs. 

Board Members and Panel Chairmen completed linal reviews of ?anel 

reports and also reviewed final draft of findings, determinations, and 

recommendations. 

3oard and Panel Members recei-$.ed a detailed trieflng on tternostatiswitch questions with emphasis upon actions of various organizatims 

during and after detankiiig operations at KSC. 

Board Members met in Executive Session and complete3 final revisions 

of Chapter 4 of the Board summary. 

Board ana Panel Members witnessed a special full-scale tank ignition 

test performed at MSC. 

Panel Chairmen completed final revisions of individual Panel reports 

and submitted copy to the Reports Editorial Office. 

Board met in Executive Session and agreed to final schedule for report printing and delivery to the Administrator on June 15, 1970. 

2-42 
Board Members met in Executive Session and completed vork on Chap 

ter j of the Board Smary Report (Findings, Determinations, and Recommendations). 

Board Members revie;red final version of Project Management Panel 

report and authorized printing as Appendix E. 

3oard Members Hedrick and Mark completed final tabulation of test 

support activities performed for the Board. 

302i-d Members revieved films of special test activities performed 

at various NASA Centers. 

JJ~ 6, 1970 

loard mt in Executive Session throughout the day and completed 

its review of Chapter 5 of its report (Findings, Determinations, 2nd 

P, e c ome nd at ions ) . 

9oard Members completed review of afialpes to be incorporated in 

Appendix F, Special Tests and Analyses. 

The Board met in Executive Session and approved plans and schedules 

for final editorial review and publication of the Board report. 

The Chaiman recessed the Board until June 1: at irh.ich time the 

Board is scheduled to reconvene in Washington, D.C., to present its 

report to t:le NASA Administrator and Deputy Ahministrator. 

2-43 
152 

CHAPTER 3 

DESCRIPTION OF APOLLO 13 SPACE VEHICLE 

AND MISSION SUMMARY 
This chapter is extracted from Mission Operation Report 

No. M-932-70, Revision 3, published by the Program and Special Reports 

Division (XP), Executive Secretariat, NASA Headquarters, Washington, 

D.C. 

Discussion in this chapter is broken into two parts. Part 1 is 

designed to acquaint the reader with the flight hardware and with 

the mission monitoring, support, and control knctions and capabilities. 

Part 2 describes the Apollo 13 mission and gives a mission sequence 

of events summary. 

3- 1 
PART 1 APOLLO/SATURN v SPACE VEHICLE 

The primary flight hardware of the Apollo Program consists of the 

Saturn V launch vehicle and Apollo spacecraft (fig. 3-11, Collectively, 

they are designated the Apollo/Saturn V soace vehicle (SV). Selected 

major systems and subsystems of the space vehicle may be smarized as 

follows. 

SATJRN V LAUNCH VEHICLE 

The Saturn V launch vehicle (LV) is designed to boost up to 

300,000 pounds into a 105-nautical mile earth orbit and to provide for 

lunar payloads of over 100,000 pounds. The Saturn V LV consists of 

three propulsive stages (S-IC, S-11, S-IVB), two interstages, and an 

instrument unit (Iu). 

S-IC Stage 

The S-IC stage (fig. 3-2) is a large cylindrical booster, 138 feet 

long and 33 feet in diameter, powered by five liquid propellant F-1 

rocket engines. These engines develop a nominal sea level thrust total 

of approximately 7,650,000 pounds. 

288,000 pounds and the total loaded stage weight is approximately 

5,031,503 pounds. The S-IC stage interfaces structurally and electrically with the S-I1 stage. It also interfaces structurally, electrically, and pneumatically with ground support equipment (GSE) through 

two umbilical service arms, three tail service masts, aqd certain 

electronic systems by antennas. The S-IC stage is instrumer,ted for 

operational measurements or signals which are transmitted by its independent telemetry system. 

The stage dry weight is approximately 

S-I1 Stage 

The S-11 stage (fig. 3-3) is a large cylindrical booster, 81.5 feet 

long and 33 feet in diameter, powered by five liquid propellant J-2 

rocket engines which develop a nominal vacuum thrust of 230,000 pounds 

each for a total of 1,150,000 pounds. Dry weight of the S-I1 stage is 

approximately 78,050 pounds. 

is 1,O75,OOO pounds. The S-IC/S-I1 interstage weighs 10,460 pounds. 

The S-I1 stage is instrumented for operational and research and development measurements which are transmitted by its independent telemetry 

system. 

the S-IC and S-IVB stages, and electric, pneumatic, and fluid interfaces 

with GSE through its umbilicals and antennas. 

The stage approximate loaded gross weight 

The S-I1 stage has structural and electrical interfaces with 

3-2 
Launch escape system 

Boost protective cover 

Command module 

Service module 

Spacecraft 

adapter 

Swc

Fi gurr 'i-~l. - Ap
156 

FLIGHT TERMINATION 

- INSTRUMENTATION 

F-1- (5) f] [ HE INSTRUMENTATION FLIGHT CONTROI. 

SERVO ACTUATO R 

RETROROCKETS 

Figure 3-2. - S-Ic S 

ENGINE FA1 RING 

tage. 

3-4 
157 

,-SYSTEMS TUNNEL / 

Figure 3-3.- S-I1 stage. 

3- 5 

47-591 0 - 70 - 11 
S-IVB Stage 

The S-IVB stage (fig. 3-4) is a large cylindrical booster 59 feet 

long and 21.6 feet in diameter, powered by one 5-2 engine. The S-IVB 

stage is capable of multiple engine starts. Engine thrust is 

203,000 pounds. This stage is also unique in that it has an attitude 

control capability independent of its main engine. Dry weight of the 

stage is 25,050 pounds. 

The interstage weight of 8100 pounds is not included in the stated 

weights. 

nals which are transmitted by its independent telemetry system. 

The launch weight of the stage is 261,700 pounds. 

The stage is instrumented for functional measurements or sigThe high performance J-2 engine as installed in the S-NB stage 

has a multiple start capability. 

to produce a thrust of 203,000 pounds during its first burn to earth 

orbit and a thrust of 178,000 pounds (mixture mass ratio of 4.5:l) 

during the first 100 seconds of translunar injection. The remaining 

translunar injection acceleration is provided at a thrust level of 

203,000 pounds (mixture mass ratio of 5.0:l): The engine valves are 

controlled by a pneumatic system powered by gaseous helium which is 

stored in a sphere inside a start bottle. An electrical control system 

that uses solid stage logic elements is used to sequence the start and 

shutdown operations of the engine. 

The S-IVB J-2 engine is scheduled 

Instrument Unit 

The Saturn V launch vehicle is guided from its launch pad into 

earth orbit primarily by navigation, guidance, and control equipment 

located in the instrument unit (IU). The instrument unit is a cylindrical structure 21.6 feet in diameter and 3 feet high installed on top of 

the S-IVB stage. The unit weighs 4310 pounds and contains measurements 

and telemetry, command communications, tracking, and emergency detection 

system components along with supporting electrical power and the environmental control system. 

APOLLG SPACECRAFT 

The Apollo spacecraft (S/C) is designed to support three men in space 

for periods up to 2 weeks, docking in space, landing on and returning 

from the lunar surface, and safely entering the earth's atmosphere. 

Apollo S/C consists of the spacecraft-to-lhl adapter (SLA) , the service 

module (SM), the command module (CM), the launch escape system (LES) , and 

the lunar module (IM). The CM and SM as a unit are referred to as the 

command and service module (CSM) . 

The 

a 

3-6 

F 
159 

FORWARD SKIRT 

10.2 

t 

IEET 

1 

AFT INTERSTAGE 

3-7 
Spacecraft-to-LM Adapter 

The SLA (fig. 3-5) is a conical structure which provides a structural 

load path between the LV and SM and also supports the LM. 

cally, the SLA smoothly encloses the irregularly shaped LM and transitions 

the space vehicle diameter from that of the upper stage of the LV to that 

of the SM. The SLA also encloses the nozzle of the SM engine and the high 

gain antenna. 

AerodynamiSpring thrusters are used to separate the Ll4 from the SLA. After 

the CSM has docked with the LM, mild charges are fired to release the 

four adapters which secure the LM in the SLA. 

spring thrusters mounted on the lower (fixed) SLA panels push against 

the LM landing gear truss assembly to separate the spacecraft from the 

launch vehicle. 

Simultaneously, four 

Service Module 

The service module (SM)(fig. 3-61 provides the main spacecraft propulsion and maneuvering capability during a mission. 

most of the spacecraft consumables (oxygen, water, propellant, and 

hydrogen) and supplements environmental, electrical power, and propulsion requirements of the CM. The SM remains attached to the CM until 

it is jettisoned just before CM atmospheric entry. 

The SM provides 

Structure.- The basic structural components are forward and aft 

(upper and lower) bulkheads, six radial beams, four sector honeycomb 

panels, four reaction control system honeycomb panels, aft heat shield, 

and a fairing. The forward and aft bulkheads cover the top and bottom 

of the SM. Radial beam trusses extending above the forward bulkhead 

support and secure the CM. The radial beams are made of solid aluminum 

alloy which has been machined and chem-milled to thicknesses varying 

between 2 inches and 0.018 inch. Wee of these beams have compression 

pads and the other three have shear-compression pads and tension ties. 

Explosive charges in the center sections of these tension ties are used 

to separate the CM from the SM. 

An aft heat shield surrounds the service propulsion engine to 

protect the SM from the engine's heat during thrusting. 

the CM and the forward bulkhead of the SM is closed off with a fairing 

which is composed of eight electrical power system radiators alternated 

with eight aluminum honeycomb panels. The sector and reaction control 

system panels are 1 inch thick and are made of aluminum honeycomb core between two aluminum face sheets. The sector panels are bolted to the 

radial beams. Radiators used to dissipate heat from the environmental 

control subsystem are bonded to the sector panels on opposite sides of 

the SM. These radiators are each about 30 square feet in area. 

The gap between 

3-8 
161 

Figure 3-5.- Spacecraft-to-m adapter. 
162 

SECTOR 2 

SECTOR 3 

SECTOR 4 

SECTOR 5 

SECTOR 6 

CENTER SECTION. SERVICE PROPULSION ENGINE AND 

HELIUM TANKS 

Figure 3-6.- Service module. 

3-10 
163 

The SM interior is divided into six sectors, or bays, and a center 

section. Sector one is currently void. It is available for installation 

of scientific or additional equipment should the need arise. Sector 

two has part of a space radiator and a reaction control system (RCS) 

engine quad (module) on its exterior panel and contains the service propulsion system (SPS) oxidizer sump tank. This tank is the larger of 

the two tanks that hold the oxidizer for the SPS engine. Sector three 

has the rest of the space radiatx and another RCS engine quad on its 

exterior panel and contains the oxidizer storage tank. This tank is 

the second of two SPS oxidizer tanks and feeds the oxidizer sump tank 

in sector two. Sector four contains most of the electrical power generating equipment. It contains three fuel cells, two cryogenic oxygen 

and two cryogenic hydrogen tanks, and a power control relay box. The 

cryogenic tanks supply oxygen to the environmental control subsystem 

and oxygen and hydrogen to the fuel cells. Sector five has part of an 

environmental control radiator and an RCS engine quad on the exterior 

panel and contains the SPS engine fuel sump tank. This tank feeds the 

engine and is also connected by feed lines to the storage tank in 

sector six. Sector six has the rest of the environmental control raditor and an RCS engine quad on its exterior and contains the SPS engine 

fuel storage tank which feeds the fuel sump tank in sector five. The 

center section contains two helium tanks and the SPS engine. The tanks 

are used to provide helium pressurant for the SPS propellant tanks. 

h.opulsion.- Main spacecraft propulsion is provided by the 

20500-pound thrust SPS. The SPS engine is a restartable, non-throttleable 

engine which uses nitrogen tetroxide (NpO4) as an oxidizer and a 50-50 

mixture of hydrazine and unsymmetrical-dirnethylhydrazine (UDMH) as fuel. 

(These propellants are hypergolic, i.e., they burn spontaneously when 

combined without need for an igniter. ) This engine is used for major 

velocity changes during the mission, such as midcourse corrections, 

lunar orbit insertion, transearth injection, and CSM aborts. The SPS 

engine responds to automatic firing commands from the guidance and 

navigation system or to commands from manual controls. The engine as- 

. sembly is gimbal-mounted to allow engine thrust-vector alignment with the 

spacecraft center of mass to preclude tumbling. Thrust-vector alignment 

control is maintained by the crew. 

about and along three axes. The SM RCS provides for maneuvering 

Additional SM systems.- In addition to the systems already described, 

the SM has communication antennas, umbilical connections, and several 

exterior mounted lights. The four antennas on the outside of the SM are 

the steerable S-band high-gain antenna, mounted on the aft bulkhead; two 

VHF omnidirectional antennas, mounted on opposite sides of the module 

near the top; and the rendezvous radar transponder antenna, mounted in 

the SM fairing. 

3-11 
164 

Seven lights are mounted in the aluminum panels of the fairing. 

Four lights (one red, one green, and two amber) are used to aid the 

astronauts in docking: one is a floodlight which can be turned on to 

give astronauts visibility during extravehicular activities, one is a 

flashing beacon used to aid in rendezvous, and one is a spotlight used 

in rendezvous from 500 feet to docking with the LM. 

SM/CM separation.- Separation of the SM from the CM occurs shortly 

before entry. The sequence of events during separation is controlled 

automatically by two redundant service module jettison controllers (SMJC) 

located on the forward bulkhead of the SM. 

Command Module 

The command module (CM) (fig. 3-7) serves as the command, control, 

and communications center for most of the mission. Supplemented by the 

SM, it provides all life support elements for three crewmen in the mission environments and for their safe return to the earth's surface. It 

is capable of attitude control about three axes and some lateral lift 

translation at high velocities in earth atmosphere. It also permits LM 

attachment, CM/LM ingress and egress, and serves as a buoyant vessel in 

open ocean. 

Structure.- The CM consists of two basic structures joined together: 

the inner structure (pressure shell) and the outer structure (heat 

shield). 

of aluminum sandwich construction consisting of a welded aluminum inner 

skin, bonded aluminum honeycomb core, and outer face sheet. The outer 

structure is basically a heat shield and is made of stainless steelbrazed honeycomb brazed between steel alloy face sheets. Parts of the 

area between the inner and outer sheets are filled with a layer of 

fibrous insulation as additional heat protection. 

The inner structure, the pressurized crew compartment, is made 

Display and controls.- The main display console (MDC) (fig. 3-8) 

has been arranged to provide for the expected duties of crew members. 

These duties fall into the categories of Commander, CM Pilot, and LJ4 

Pilot, occupying the left, center, and right couches, respectively. The 

CM Pilot also acts as the principal navigator. All controls have been 

designed so they can be operated by astronauts wearing gloves. The con- trols are predominantly of four basic types: toggle switches, rotary 

switches with click-stops, thumb-wheels, and push buttons. Critical 

switches are guarded so that they cannot be thrown inadvertently. In 

addition, some critical controls have locks that must be released before 

they can be operated. 

3-12 
165 

/ ROLL ENGINES 

(TYPICAL) 8 BAND ANTENNA 

-2 -y -x 

LEFT HAND COWBINED TUNNEL hATCH 

FORVIARD COMPARTMENT, /FOR,VARD EOUIPMENT WY / 

COMPARTMENT 

\ 'LEFT HAND EQUIPMENT 84Y RIGHT HAND EOUIPMENT UY 

AFT COMPARTMENT AFT COMPARTMEWT 

Figure 3-7.- Ccmmanu module, 

3 -13 
166 

rService 

)pu Is ion 

Cryogenics 7 

contro1\f9 

and warning 

-SCS power panel Environmental control2 

n 

Launch vehicle emergency detection 

Flight attitude 

Mission sequence 

Velocity change monitor 

Entry monitor 

Propellant gauging 

Environment control 

Communications control 

Power distribution 

Caution and warning 

7, 

r igire 5-8. - CEI main dispizy console. 

3-14 
167 

Flight controls are located on the left center and left side of the 

MDC, opposite the Commander. These include controls for such subsystems 

as stabilization and control, propulsion, crew safety, earth landing, 

and emergency detection. One of two guidance and navigation computer 

panels also is located here, as are velocity, attitude, and altitude 

indicators. 

The CM Pilot faces the center of the console, and thus can reach 

many of the flight controls, as well as the system controls on the right 

side of the console. Displays and controls directly opposite him include 

reaction control, propellant management, caution and warning, environmental control, and cryogenic storage systems. The rotation and translation controllers used for attitude, thrust vector, and translation 

maneuvers are located on the arms of two crew couches. In addition, a 

rotation controller can be mounted at the navigation position in the 

lower equipment bay. 

Critical conditions of most spacecraft systems are monitored by a 

caution and warning system. A malfunction or out-of-tolerance condition 

results in illumination of a status light that identifies the abnormal- I 

ity. It also activates the master alarm circuit, which illuminates two 

master alarm lights on the MDC and one in the lower equipment bay and 

sends an alarm tone to the astronauts' headsets. The master alarm 

lights and tone continue until a crewman resets the master alarm circuit. 

This can be done before the crewmen deal with the problem indicated. The 

caution and warning system also contains equipment to sense its own 

malfunctions. 

Lunar Module 

The lunar module (m) (fig. 3-9) is designed to transport two men 

safely from the CSM, in lunar orbit, to the lunar surface, and return 

them to the orbiting CSM. The LM provides operational capabilities such 

as communications, telemetry, environmental support, transportation of 

scientific equipment to the lunar surface, and returning surface samples 

with the crew to the CSM. 

The lunar module consists of two stages: the ascent stage and the 

Separable umbilicals and hardline connections provide subsystem 

The LM is designed to operate for 48 hours 

descent stage. The stages are attached at four fittings by explosive 

bolts. 

continuity to operate both stages as a single unit until separate ascent 

stage operation is desired. 

after separation from the CSM, with a maximum lunar stay time of 44 hours. 

Table 3-1 is a weight summary of the Apollo/Saturn 5 space vehicle for 

the Apollo 13 mission. 

3-15 
168 

Overhead 

S-band hatch 

RCS thrust chamber 

Figure 3-9.- Lunar module. 

3-16 
169 

TABLE 3-1.- APOLLO 13 WEIGHT SUMMARY (WEIGHT IN POUNDS) 

Stage/module 

s-IC 

s-IC/S-I1 

interstage 

S-I1 stage 

S-II/S-IvB 

i n t e r s t age 

S-IVB stage 

Instrument unit 

Spacecraft-LM 

adapter 

Lunar module 

Service module 

Command module 

Launch escape 

system 

Inert weight 

288000 

11464 

78050 

8100 

25050 

4482 

Tot a1 

expendables 

4746870 

--- 

996960 

--- 

236671 

--- 

Total weigh 

5034870 

11464 

1075010 

8100 

261721 

4482 

Launch vehicle at ignition 6,395,647 

4044 

9915 

10532 

12572 

9012 

4044 

33483 

51099 

12572 

9012 

Final 

3eparation 

weight 

--- 

*33941 

**14076 

**11269 

Landing) 

--- 

* CSM/LM separation 

** CM/SM separation 

3-17 
170 

TABLE 3-1.- APOLLO 13 WEIGHT SUMMARY (WEIGHT IN POUNDS) - Concluded 

Final 

weight 

Stage/module I Inert weight I eq~~~~~les 1 Total weight1 separation 

Spacecraft at ignition 110,210 

Space vehicle at ignition 

S-IC thrust buildup 

Space vehicle at lift-off 

Space vehicle at orbit insertion 

6505857 

(-)a4598 

6421259 

299998 

Main propulsion.- Main propulsion is provided by the descent propulsion system (DPS) and tfie ascent propulsion system (APS). Each 

system is wholly independent of the other. The DPS provides the thrust 

to control descent to the lunar surface. The APS can provide the thrust 

for ascent from the lunar surface. In case of mission abort, the APS 

and/or DPS can place the LM into a rendezvous trajectory with the CSM 

from any point in the descent trajectory. The choice of engine to be 

used depends on the cause for abort, on how long the descent engine 

has been operating, and on the quantity of propellant remaining in the 

descent stage. Both propulsion systems use identical hypergolic propellants. The fuel is a 50-50 mixture of hydrazine and unsymmetricaldimethylhydrazine and the oxidizer is nitrogen tetroxide. Gaseous 

helium pressurizes the propellant feed systems. Helium storage in the 

DPS is at cryogenic temperatures in the super-critical state and in the 

APS it is gaseous at ambient temperatures. 

Ullage for propellant settling is required prior to descent engine 

start and is provided by the +X axis reaction engines. 

engine is gimbaled, throttleable, and restartable. The engine can be 

throttled from 1050 pounds of thrust to 6300 pounds. 

above this value automatically'produce full thrust to reduce combustion 

chamber erosion. Nominal full thrust is 9870 pounds. Gimbal trim of 

the engine compensates for a changing center of gravity of the vehicle 

and is automatically accomplished by either the primary guidance and 

navigation system (PGNS) or the abort guidance system (AGS). Automatic 

throttle and on/off control is available in the PGNS mode of operation. 

The descent 

Throttle positions 

3-18 
171 

The AGS commands on/off operation but has no automatic throttle control 

capability. Manual control capability of engine firing functions has 

been provided. Manual thrust control override may, at any time, command more thrust than the level commanded by the LM guidance computer 

(LGC). 

develops 3500 pounds of thrust, sufficient to abort the lunar descent 

or to launch the ascent stage from the lunar surface and place it in 

the desired lunar orbit. Control modes are similar to those described 

for the descent engine. The APS propellant is contained in two spherical titanium tanks, one for oxidizer and the other for fuel. Each tank 

has a volume of 36 cubic feet. Total fuel weight is 2008 pounds, of 

which 71 pounds are unusable. Oxidizer weight is 3170 pounds, of which 

92 pounds are unusable. The APS has a limit of 35 starts, must have a 

propellant bulk temperature between 50' F and 90' F prior to start, 

must not exceed 460 seconds of burn time, and has a system life of 

24 hours after pressurization. 

The ascent engine is a fixed, non-throttleable engine. The engine 

Electrical power system.- The electrical power system (EPS) contains six batteries which supply the electrical power requirements of 

the LM during undocked mission phases. Four batteries are located in 

the descent stage and two in the ascent stage. Batteries for the 

explosive devices system are not included in this system description. 

Postlaunch LM power is supplied by the descent stage batteries until 

the LM and CSM are docked. While docked, the CSM supplies electrical 

power to the LM up to 296 watts (peak). 

the two ascent stage batteries are paralleled with the descent stage 

batteries for additional power assurance. The descent stage batteries 

are utilized for LM lunar surface operations and checkout. The ascent 

stage batteries are brought on the line just before ascent phase 

staging. p11 batteries and busses may be individually monitored for 

load, voltage, and failure. Several isolation and combination modes 

are provided. 

During the lunar descent phase, 

Two inverters, each capable of supplying full load, convert the 

dc to ac for 115-volt, 400-hertz supply. 

by the following busses: LM Pilot's dc bus, Commander's dc bus, and ac 

busses A and B. 

Electrical power is distributed 

The four descent stage silver-zinc batteries are identical and have 

a 400 ampere-hour capacity at 28 volts. 

have a constant voltage at various states of charge/load levels, "high" 

and "low" voltage taps are provided for selection. 

tap is selected to initiate use of a fully charged battery. Cross-tie 

circuits in the busses facilitate an even discharge of the batteries 

regardless of distribution combinations. The two silver-zinc ascent 

stage batteries are identical to each other and have a 296 ampere-hour 

Because the batteries do not 

The "low voltage" 

3-19 
172 

capacity at 28 volts. The ascent stage batteries are normally connected 

in parallel for even discharge. Because of design load characteristics, 

the ascent stage batteries do not have and do not require high and low 

voltage taps. 

Nominal voltage for ascent stage and descent stage batteries is 

30.0 volts. Reverse current relays for battery failure are one of many 

components designed into the EPS to enhance EPS reliability. Cooling 

of the batteries is provided by the environmental control system cold 

rail heat sinks. Available ascent electrical energy is 17.8 kilowatt 

hours at a Iriaximum drain of 50 amps per battery and descent energy is 

46.9 kilowatt hours at a maximum drain of 25 anips per battery. 

MISSION MONITORING, SUPPORT, AND CONTROL 

Mission execution involves the following functions: prelaunch 

checkout and launch operations; tracking the space vehicle to determine 

its present and future positions; securing information on the status of 

the flight crew and space vehicle systems (via telemetry); evaluation 

of telemetry information; commanding the space vehicle by transmitting 

real-time and updata commands to the onboard computer; and voice communication between flight and ground crews. 

These functions require the use of a facility to assemble and 

launch the space vehicle (see Launch Complex), a central flight control 

facility, a network of remote stations located strategically around the 

world, a method of rapidly transmitting and receiving information 

between the space vehicle and the central flight control facility, and 

a real-time data display system in which the data are made available 

and presented in usable form at essentially the same time that the data 

event occurred. 

The flight crew and the following organizations and facilities 

participate in mission control operations: 

a. Mission Control Center (MCC), Manned Spacecraft Center (MSC), 

Houston, Texas. The MCC contains the communication, computer display, 

and command systems to enable the flight controllers to effectively 

monitor and control the space vehicle. 

b. Kennedy Space Center (KSC), Cape Kennedy, Florida. The space 

vehicle is launched from KSC and controlled from the Launch Control 

Center (LCC). 

at the Central Instrumentation Facility (CIF) at KSC from the launch 

pads, CIF receivers, Merritt Island Launch Area (MILA), and the downrange Air Force Eastern Test Range (PEETR) stations. These data are 

Prelaunch, launch, and powered flight data are collected 

3-20 
173 

transmitted to MCC via the Apollo Launch Data System (ALDS). 

located at KSC (AFETR) is the Impact Predictor (IP), for range safety 

purposes. 

Also 

c. Goddard Space Flight Center (GSFC), Greenbelt, Maryland. GSFC 

manages and operates the Manned Space Flight Network (MSFN) and the 

NASA communications (NASCOM) network. During flight, the MSFN is 

under the operational control of the MCC. 

d. George C. Marshall Space Flight Center (MSFC) , Huntsville, 

Alabama. MSFC, by means of the Launch Information Exchange Facility 

(LIEF) and the Huntsville Operations Support Center (HOSC) provides 

launch vehicle systems real-time support to KSC and MCC for preflight, 

launch, and flight operations. 

A block diagram of the basic flight control interfaces is shown 

in figure 3-10. 

Vehicle Flight Control Capability 

Flight operations are controlled from the MCC. The MCC has two 

flight control rooms, but only one control room is used per mission. 

Each control room, called a Mission Operations Control Room (MOCR) , is 

capable of controlling individual Staff Support Rooms (SSR's) located 

adjacent to the MOCR. The SSR's are manned by flight control specialists who provide detailed support to the MOCR. Figure 3-11 outlines 

the organization of the MCC for flight control and briefly describes 

key responsibilities. Information flow within the MOCR is shown in 

figure 3-12. 

The consoles within the MOCR and SSR's permit the necessary interface between the flight controllers and the spacecraft. The displays 

and controls on these consoles and other group displays provide the 

capability to monitor and evaluate data concerning the mission and, 

based on these evaluations, to recommend or take appropriate action on 

matters concerning the flight crew and spacecraft. 

Problems concerning crew safety and mission success are identified 

to flight control personnel in the following ways: 

a. Flight crew observations 

b. Flight controller real-time observations 

c. Review of telemetry data received from tape recorder playback 

d. Trend analysis of actual and predicted values 

3-21 

ii-591 0 - 70 - 12 
174 

e. Review of collected data by systems specialists 

f. Correlation and comparison with previous mission data 

g. Analysis of recorded data from launch complex testing 

3-22 
175 

3-2 3 
176 

OVERALL CONDUCT OF 

RECOVERYANUOTHER 

TO PUBLIC MISSION SUPPORT 

I DECISIONSIACTIONS ON SPACE 

AN0 MCCIMSFN OPERATIONS 

I VEHICLE SVSTEMSlDINAMlCS 

MISSION COMMAND 

ANOCONTROL GROUP I SVSTEMS OPERATIONS 

GROUP 

1-1 I-+ 

FLIGHT DYNAMICS 

GROUP 

MONITORS PRELAUNCH CHECKOU 

PONERED FLIGHT EVENTS AN0 

TRAECTORIES REENTRY EVENTS 

MAINTAINS UPDATE0 ABORT 

ANDREENTRYPLAN UPDATES 

MONITORS GUIOANCE 

FUNCTIONS WRING POWERED 

FLIGHT AND PREMANEUVER 

PREPARATION 

Figure 3-11.- Mission Control Center organization. 

3-2 4 
177 

[MIjSiON] DIRECTOR 

VEHICLE 

VEHICLE 

SYSTEMS 

FLIGHT FLIGHT 

DIRECTOR 

FLIGHT - 

MI:sl:N 

PROCEDURE 

I MISSYN PROCEDURE STATUS I , CF~:~HT & 

STATUS COMMUNICATOR 

0 AND P 

OFFICER 

Figure 3-12.- Information flow within the 

Mission Operations Control Room. 

3-25 
178 

PART 2. APOLLO 13 MISSION DESCRIPTION 

PRIMARY MISSION OBJECTIVES 

The primary mission objectives were as follows: 

Perform selenological inspection, survey, and sampling of materials 

in a preselected region of the Fra Macro Formation. 

Deploy and activate an Apollo Lunar Surface Experiments Package 

( ALSEP ) . 

Develop man's capability to work in the lunar environment. 

Obtain photographs of candidate exploration sites. 

Table 3-11 lists the Apollo 13 mission sequence of major events and 

the time of occurrence in ground elapsed time. 

TABLE 3-11. - APOLLO 13 MISSION SEQUENCE OF EVENTS 

Event 

Range zero (02:13:00.0 p.m. e.s.t., April 11) 

Earth parking orbit insertion 

Second S-IVB ignition 

Trans lunar injection 

CSM/S-IVB separation 

Spacecraft ejection from S-IVB 

S-IVB APS evasive maneuver 

S-IVB APS maneuver for lunar impact 

Midcourse correction - 2 (hybrid transfer) 

Cryogenic oxygen tank anomaly 

Midcourse correction - 4 

S-IVB lunar impact 

Pericynthion plus 2-hour maneuver 

Midcourse correction - 5 

Midcourse correction - 7 

Service module jettison 

Lunar module jettison 

Entry interface 

Landina 

Ground elapsed time 

(hr:min:sec) 

00 :oo :oo 

00 : 12 : 40 

02 : 35 :46 

02: 41:47 

03 :06 : 39 

04: 01 :03 

Oh :18: 01 

30: 40 : 50 

61:29 : 43 

7736: 40 

79 :27 : 39 

105 : 18 : 32 

137:39:49 

138 :02 :06 

141:30:02 

142: 40 : 47 

142 : 54 : 41 

05 : 59 : 59 

55:54:53 

3-26 
179 

Launch and Earth Parking Orbit 

Apollo 13 was successfilly launched on schedule from Launch Complex 

39A, Kennedy Space Center, Florida, at 2:13 p.m. e.s.t., April 11, 1970. 

The launch vehicle stages inserted the S-IVB/instrument unit (IU)/ 

spacecraft combination into an earth parking orbit with an apogee of 

100.2 nautical miles (n. mi.) and a perigee of 98.0 n. mi. (100-n. mi. 

circular planned). 

S-I1 stage cut off about 132 seconds early, causing the remaining four 

engines to burn approximately 34 seconds longer than predicted. Space 

vehicle velocity after S-I1 boost was 223 feet per second (fps) lower 

than planned. 

imately 9 seconds longer than predicted with cutoff velocity within 

about 1.2 fps of planned. Total launch vehicle burn time was about 

44 seconds longer than predicted. 

meeting translunar injection (TLI) cutoff conditions existed with remaining S-IVB propellants. 

During second stage boost, the center engine of the 

As a result, the S-IVB orbital insertion burn was approxA greater than 3-sigma probability of 

After orbital insertion, all launch vehicle and spacecraft systems 

were verified and preparation was made for translunar injection (TLI). 

Onboard television was initiated at 01:35 ground elapsed time (g.e.t.) 

for about 5.5 minutes. The second S-IVB burn was initiated on schedule 

for TLI. 

ditions were nominal for a free-return circumlunar trajectory. 

All major systems operated satisfactorily and all end conTranslunar Coast 

The CSM separated from the LM/IU/S-IVB at about 03:07 g.e.t. Onboard television was then initiated for about 72 minutes and clearly 

showed CSM "hard docking," ejection of the CSM/LM from the S-IVB at 

about 04:Ol g.e.t., and the S-IVB auxiliary propulsion system (APS) 

evasive maneuver as well as spacecraft interior and exterior scenes. 

!be SM RCS propellant usage for the separation, transposition, docking, 

and ejection was nominal. All launch vehicle safing activities were 

performed as scheduled. 

The S-IVE APS evasive maneuver by an 8-second APS Ullage burn was 

initiated at 04:18 g.e.t. and was successfully completed. The liquid 

oxygen dump was initiated at 04:39 g.e.t. and was also successfully 

accomplished. The first S-IVE APS burn for lunar target point impact 

was initiated at 06:oo g.e.t. 

ducing a differential velocity of approximately 28 ms. Tracking information available at 08:OO g.e.t. indicated that the S-IVB/IU would impact 

at 6'53' S., 30'53' W. versus the targeted 3' S., 30' W. Therefore, the 

second S-IVB APS (trim) burn was not required. 

sure dropped in the IU ST-124-M3 inertial platform at 18:25 g.e.t. and 

the S-IVB/IU no longer had attitude control but began tumbling slowly. 

The burn duration was 217 seconds, proThe gaseous nitrogen pres3-27 
At approximately 19:17 g.e.t., a step input in tracking data indicated a 

velocity increase of approximately 4 to 5 fps. No conclusions have been 

reached on the reason for this increase. The velocity change altered 

the lunar impact point closer to the target. 

lunar surface at 77:56:40 g.e.t. (08:09:40 p.m. e.s.t. April 14) at 

2.4' S., 27.9' W., and the seismometer deployed during the Apollo 12 

mission successfully detected the impact. 

125 n. mi. from the seismometer. The actual impact point was 74 n. mi. 

from the seismometer, well within the desired 189-n. mi. (350-km) radius. 

The S-IVB/IU impacted the 

The targeted impact point was 

The accuracy of the TLI maneuver was such that spacecraft midcourse 

correction No. 1 (MCC-l), scheduled for 11:hl g.e.t., was not required. 

MCC-2 was performed as planned at 30:41 g.e.t. and resulted in placing 

the spacecraft on the desired, non-free-return circumlunar trajectory 

with a predicted closest approach to the moon on 62 n. mi. 

parameters were normal. The accuracy of MCC-3 was such that MCC-3, 

scheduled for 55:26 g.e.t., was not performed. 

coverage of the preparations and performance of MCC-2 was received for 

49 minutes beginning at 30:13 g.e.t. 

All SPS burn 

Good quality television 

At approximately 55:55 g.e.t. (10:08 p.m. e.s.t.), the crew reported an undervoltage alarm on the CSM main bus B. Pressure was rapidly lost in SM oxygen tank no. 2 and fuel cells 1 and 3 current dropped 

to zero due to loss of their oxygen supply. A decision was made to 

abort the mission. The increased load on fuel cell 2 and decaying pressure in the remaining oxygen tank led to the decision to activate the 

LM, power down the CSM, and use the LM systems for life support. 

At 61:30 g.e.t., a 38-fps midcourse maneuver (MCC-4) was performed 

by the LM DPS to place the spacecraft in a free-return trajectory on 

which the CM would nominally land in the Indian Ocean south of Mauritius 

at approximately 152:OO g.e.t. 

Transearth Coast 

At pericynthion plus 2 hours (79:28 g.e.t.), a LM DPS maneuver was 

The 263.4-second burn produced a differential velocity of 860.5 

performed to shorten the return trip time and move the earth landing 

point. 

fps and resulted in an initial predicted earth landing point in the midPacific Ocean at 142:53 g.e.t. 

up and the primary system was used for this maneuver. 

maneuver, passive thermal control was established and the LM was powered 

down to conserve consumables; only the LM environmental control system 

(ECS) and communications and telemetry systems were kept powered up. 

Both LM guidance systems were powered 

Following the 

The LM DPS was used to perform MCC-5 at 105:19 g.e.t. The 15-second 

burn (at 10-percent throttle) produced a velocity change of about 7.8 fps 

3-28 
and successfully raised the entry flight path angle to -6.52'. 

The CSM was partially powered up for a check of the thermal conditions of the CM with first reported receipt of S-band signal at 101:53 

g.e.t. Thermal conditions on all CSM systems observed appeared to be in 

order for entry. 

Due to the unusual spacecraft configuration, new procedures leading 

to entry were developed and verified in ground-based simulations. The 

resulting timeline called for a final midcourse correction (MCC-7) at 

entry interface (EI) -5 hours, jettison of the SM at EI -4.5 hours, then 

jettison of the LM at EI -1 hour prior to a normal atmospheric entry by 

the CM. 

MCC-7 was successfully accomplished at 137:40 g.e.t. The 22.4-second 

LM RCS maneuver resulted in a predicted entry flight path angle of -6.49'. 

The SM was jettisoned at 138:02 g.e.t. 

the SM and reported that an entire panel was missing near the S-band highgain antenna and a great deal of debris was hanging out. The CM was powered up and then the LM was jettisoned at 141:30 g.e.t. The EI at 40,000 

feet was reached at 142:41 g.e.t. 

The crew viewed and photographed 

Entry and Recovery 

Weather in the prime recovery area was as follows: broken stratus 

clouds at 2000 feet; visibility 10 miles; 6-knot ENE winds; and wave 

height 1 to 2 feet. Drogue and main parachutes deployed normally. 

Visual contact with the spacecraft was reported at 142:50 g.e.t. Landing 

occurred at 142:54:41 g.e.t. (01:07:41 p.m. e.s.t., April 17). The landing point was in the mid-Pacific Ocean, approximately 21'40' s., 165'22' W. 

The CM landed in the stable 1 position about 3.5 n. mi. from the prime 

recovery ship, USS IWO JIM. 

copter, was safe aboard the ship at 1:53 p.m. e.s.t., less than an hour 

after landing. 

The crew, picked up by a recovery heli3-29 
182 

CHAPTER 4 

mvnw AND ANALYSIS OF APOLLO 13 ACCIDENT 
183 

PART 1. INTRODUCTION 

It became clear in the course of the Board's review that the accident during the Apollo 13 mission was initiated in the service module 

cryogenic oxygen tank no. 2. Therefore, the following analysis centers 

on that tank and its history. In addition, the recovery steps taken in 

the period beginning with the accident and continuing to reentry are 

discussed. 

Two oxygen tanks essentially identical to oxygen tank no. 2 on 

Apollo 13, and two hydrogen tanks of similar design, operated satisfactorily on several unmanned Apollo flights and on the Apollo 7, 8, 9, 10, 

11, and 12 manned missions. With this in mind, the Board placed particular emphasis on each difference in the history of oxygen tank no. 2 from 

the history of the earlier tanks, in addition to reviewing the design, 

assembly, and test history. 

4-1 
PART 2. OXYGEN TANK NO. 2 HISTORY 

DESIGN 

On February 26, 1966, the North American Aviation Corporation, now 

North American Rockwell (NR), prime contractor for the Apollo command 

and service modules (CSM), 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. 

As the simplified draving in figure 4-1 indicates, each oxygen tank 

has an outer shell and an inner shell, arranged to provide a vacuum 

space to reduce heat leak, and 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 thermostatically protected heater coils 

and two small fans driven by 1800 rpm motors to stir the tar& contents. 

The other, called the quantity probe, consists of an upper section which 

supports a cylindrical capacitmce gage used to measure electrically the 

quantity of fluid in the tank. The inner cylinder of this probe serves 

both as a fill and drain tube and as one plate of the capacitance gage. 

In addition, a temperature sensor is mounted on the outside of the quantity probe near the head. Wiring for the gage, the temperature sensor, 

the fan motors, and the heaters passes through the head of the quantity 

probe to a conduit in the dome. From there the wiring runs to a connecter which ties it electrically to the appropriate external circuits 

in the CSM. The routing of wirfng and lines from the tank through the 

dome is shown in figure 4-2. 

As shown in figure 4-2, the fill line from the exterior of the SM 

enters the oxygen tank and connects to the inner cylinder of the capacitance gage through a coupling of two Teflon adapters or sleeves and a 

short length of inconel tubing. The dimensions and tolerances selected 

are such that if "worst case" variations in an actual system were to 

occur, the coupling might pot reach from the fill line to the gage cylinder (fig. 4-3). 

fit would result. 

Thus, the variations might be such that a very loose 

The supply line from the tank leads from the head of the quantity 

probe to the dome and thence, after passing around the tank between the 

inner and outer shells, exits through the dome to supply oxygen to the 

fuel cells in the service module (SM) and the environmental control 

system (ECS) in the comand module (CM). The supply line also connects 

4-2 
185 

Blowout disc Closeout cap 

Supply 

line - 

To fuel 

cell/ ECS i 

Pressure 

C 

transduce 

Figure 4-1.- Oqgen tank no. 2 internal components. 

4-3 
186 

Figure 4-2.- Oxygen tank wiring and lines. 

4- 4 
L8T 
188 

I 

I 

to a relief valve. Under normal conditions, pressure in the tank is 

measured by a pressure gage in the supply line and a pressure switch 

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. 

The oxygen tank is designed for a capacity of 320 pounds of supercritical oxygen at pressures ranging between 865 to 935 pounds per 

square inch absolute (psia). The tank is initially filled with liquid 

oxygen at -29'7' F and operates over the range from -340' F to +80° F. 

Tne temi "supercritical" means that the oxygen is maintained at a temperature and pressure which assures that it is a homogeneous, single-phase 

fluid. 

The burst pressure of the orygen tank is about 2200 psi at -150' F, 

over twice the normal operating pressure at that temperature. The relief 

valve is designed to relieve pressure in the oxygen tank overboard at a 

pressure 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. 

The approximate amounts of principal materials within the oxygen 

tank are set forth in table 4-1. 

I 

I 

TABLE 4-1.- MATERIALS WITHIN OXYGEN T.4NK 

Aluminum (all forms) 1 0.6 

Material 

I 20,500 

Approximate 

quantity, lb 

Inconel alloys 

Available 

energy, Btu 

1.7 2,900 

Teflon-wire insulation 

sleeving and solid 

1.1 2,400 

Stainless steel I 2.4 15,000 

Two oxygen tanks are mounted on a shelf in bay 4 of the SM, as 

shown in figure 4-4. Figures 4-5 through 4-8 are photographs of portions 

4-6 
189 

t .A 

Figure 4-4.- Arrangement of fuel cells and cryogenic systems in bay 4. 

4- 7 

47-591 0 - I0 - 13 
190 

Figure 4-5.- Fuel cells shelf. 

4-9/4-1@ 
191 

Figure 4-6.- Oxygen tank shelf 

4-11/4- 12 
192 

4-7.- 

Figure X4-B X Hydrogen tank shelf. 
193 

4-8. - Figure&S)XK Inside view of panel covering bay 4. 

4-15/4-16 
of the Apollo 13 service module (SM 109) at the North American Rockwell 

plant prior to shipment to KSC. 

with fuel cell 1 on the right, fuel cell 3 on the left, and fuel cell 2 

behind cells 1 and 3. 

lower left. Figure 4-6 shows the oxygen tank shelf, with oxygen tank 

no. 2 at left center. 

hydrogen tank no. 1 on top and hydrogen tank no. 2 below. The bottom 

of the oxygen shelf shows some of the oxygen system instrumentation and 

wiring, largely covered by insulation. Figure 4-8 is a photograph of 

the bay 4 panel, which was missing from the service module after the 

accident. 

Figure 4-5 shows the fuel cell shelf, 

The top of oxygen tank no. 2 can be seen at the 

Figure 4-7 shows the hydrogen tank shelf with 

A more detailed description of the oxygen tank design is contained 

in Appendix D to this report. 

MANUFACTURE 

The manufacture of oxygen tank no. 2 began in 1966. Under subcontracts with Beech, the inner shell of the tank was manufactured by the 

Airite Products Division of Electrada Corporation; the quantity probe 

was made by Simonds Precision Products, Inc. ; and the fans and fan 

motors were produced by Globe Industries, Inc. 

The Beech serial number assigned to the oxygen tank no. 2 flown 

in the Apollo 13 was 10024XTAOOO8. 

tank built. 

by Beech. 

It was the eighth Block I1 oxygen 

Twenty-eight Block I oxygen tanks had previously been built 

The design of the oxygen tank is such that once the upper and lower 

halves of the inner and outer shells are assembled and welded, the 

heater assembly must be inserted in the tank, moved to one side, and 

bolted in place. Then the quantity probe is inserted into the tank and 

the heater assembly wires (to the heaters, the thermostats, and the fan 

motors) must be pulled through the head of the quantity probe and the 

32-inch coiled conduit in the dome. Thus, the design 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 detected before the 

tank is capped off and welded closed. 

Several minor manufacturing flaws were discovered in oxygen tank 

no. 2 in the course of testing. A porosity in a weld on the lower half 

of the outer shell necessitated grinding and rewelding. 

also required when it was determined that incorrect welding wire had 

been inadvertently used for a small weld on a vacuum pump mounted on 

Rewelding was 

4-17 
195 

the outside of the tank dome. The upper fan motor originally installed 

was noisy and drew excessive current. The tank was disassembled and the 

heater assembly, fans, and heaters were replaced with a new assembly 

and new fans. The tank was then assembled and sealed for the second 

time, and the space between the inner and outer shells was pumped down 

over a 28-day period to create the necessary vacuum. 

TANK TESTS AT BEECH 

Acceptance testing of oxygen tank no. 2 at Beech included extensive 

dielectric, insulation, and functional tests of heaters, fans, and vacion pumps. The tank was then leak tested at 500 psi and proof tested 

at 1335 psi with helium. 

Arter the helium proof test, the tank was filled with liquid oxygen 

and pressurized to a proof pressure of 1335 psi by use of the tank 

heaters powered by 65 V ac. Extensive heat-leak tests were run at 

900 psi for 25 to 30 hours over a range of ambient conditions and outflow rates. At the conclusion of the heat-leak tests, about 100 pounds 

of oxygen remained in the tank. About three-fourths of this was released 

by venting the tank at a controlled rate through the supply line to 

about 20 psi. The tank was then emptied by applying warm gas at about 

30 psi to the vent line to force the liquid oxygen (LOX) in the tank out 

the fill line (see fig. 4-2). No difficulties were recorded in this 

detanking operation. 

The acceptance test indicated that the rate of heat leak into the 

tank was higher than permitted by the specifications. 

working, the rate improved, but was still somewhat higher than specified. 

The tank was accepted with a formal waiver of this condit’ion. Several 

other minor discrepancies were also accepted. These included oversized 

holes in the support for the electrical plug in the tank dome, and an 

oversized rivet hole in the heater assembly just above the lower fan. 

None of these items were serious, and the tank was accepted, filled with 

helium at 5 psi, and shipped to NR on Mey 3, 1967. 

After some reASSFKBLY AND TEST AT NORTH AMERICAN ROCKWELL 

The assembly of oxygen shelf serial number 0632AAG3277, with Beech 

oxygen tank serial nmber 10024XTAOOO9 as oxygen tank no. 1 and serial 

number 1002hXTA0008 as oxygen tank no. 2, was completed on March 11, 1968. 

The shelf was to be installed in SM 106 for flight in the Apollo 10 

mission. 

4-18 
196 

Beginning on April 27, the assembled oxygen shelf undewent 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 no. 2, and therefore no rework of oxygen tank no. 2 was 

required. None of the oxygen tank testing at NR requires use of LOX 

in the tanks. 

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. No anomalies were noted. 

Due to electromagnetic interference problems with the vac-ion 

pumps on cryogenic tank domes in earlier Apollo spacecraft, a modification was introduced and a decision was made to replace the complete 

oxygen shelf in SM 106. 

prepared for installation in SM 106. 

shelf was removed from SM 106 for the required modification and installation in a later spacecraft. 

An oxygen shelf with approved modifications was On October 21, 1968, the oxygen 

The oxygen shelf was removed in the manner sham in figure 4-9. 

After various lines and wires were disconnected and bolts which hold 

the shelf in the SM were removed, a fixture suspended from a crane was 

placed under the shelf and used to lift the shelf and extract it from 

bay 4. One shelf bolt was mistakenly left in place during the initial 

attempt to remove the shelf; and as a consequence, after the front of 

the shelf was raised about 2 inches, the fixture broke, allowing the 

shelf to drop back into place. Photographs of the underside of the 

fuel cell shelf in SM 1.06 indicate that the closeout cap on the dome 

of oxygen tank no. 2 may have struck the underside of that shelf during 

this incident. At the time, however, 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 2 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 oxygen shelf was removed without further difficulty. Following 

removal, the oxygen shelf was retested to check shelf integrity, including proof-pressure tests, leak tests, and f’unctional tests of 

pressure transducers and switches, thermal Switches, and vac-ion pumps. 

No cryogenic testing was conducted. 

problem. 

internal malfunctions of most types, but would not disclose fill line 

leakage within oxygen tank no. 2. 

ducted during this investigation, however, have indicated that the 

forces experienced by the shelf were probably close to those originslly 

Visual inspection revealed no 

These tests would have disclosed external leakage or serious 

Further calculations and tests con4-19 
197 

QII 

/ 

---A \ 

n in 

4-20 
calculated assuming a 2-inch drop only. 

from this incident, therefore, is now considered to be rather low, 

although it is possible that a loosely fitting fill tube could have 

been displaced by the event. 

The probability of tank damage 

The shelf passed these tests and was installed in SM 109 on 

November 22, 1968. 

were repeated in SM 109 in late December and early January, with no 

significant problems, and SM 109 was shipped to Kennedy Space Center 

(KSC) in June of 1969 for further testing, assembly on the launch 

vehicle, and launch. 

The shelf tests accomplished earlier in SM 106 

TESTING AT KSC 

At the Kennedy Space Center the CM and the SM were mated, checked, 

assembled on the Saturn V launch vehicle, and the total vehicle was moved to the launch pad. 

me countdown demonstration test (CDDT) began on March 16, 1970. 

Up to this point, nothing unusual about oxygen tank no. 2 had been 

noted during the extensive testing at KSC. 

evacuated to 5mm Hg followed by an oxygen pressure of about 80 psi. 

After the cooling of the fuel cells, cryogenic oxygen loading and tank 

pressurization to 331 psi were completed without abnormalities. At the 

time during CDDT when the oxygen tanks are normally partially emptied 

to about 50 percent of capacity, oxygen tank no. 1 behaved normally, 

but oxygen tank no. 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. An Interim Discrepancy Report was written and 

transferred to a Ground Support Equipment (GSE) Discrepancy Report, 

since a GSE filter was suspected. 

The oxygen tanks were 

On Friday, March 27, 1970, detanking operations were resumed, after 

discussions of the problem had been held with KSC, MSC, NR, and Beech 

personnel participating, either personally or by telephone. 

step, oxygen tank no. 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. 

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 sleeves and tube, 

that such a leak would allow the gaseous oxygen (GOX) being supplied 

to the vent line to leak directly to the fill line without forcing any 

As a first 

Further discussions between KSC, MSC, NR, 

Referring to figure 4-2, it will be noted 

4-21 
199 

significant mount of LOX out of the tank. 

ancy report against the spacecraft system was written. 

At this point, a discrepA "normal" detanking procedure was then conducted on both oxygen 

tanks, pressurizing through the vent line and opening the fill lines. 

Tank no. 1 emptied in a few minutes. Tank no. 2 did not. Additional 

attempts were made with higher pressures without effect, and a decision 

was made to try to "boil off" the remaining owgen in tank no. 2 by 

use of the tank heaters, 

GSE power supply, and, about 1-1/2 hours later, the fans were turned 

on to add more heat and mixing. 

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 was pressurized to about 300 psi, held for a few 

minutes, and then vented through the fill line. Tne first cycle 

produced a 7-percent quantity decrease, and the process was continued, 

with the tank emptied after five pressure/vent cycles. 

heaters were turned off after about 8 hours of heater operation. 

The heaters were energized with the 65 V dc. 

After 6 hours of heater operation, 

The fans and 

Suspecting the loosely fitting fill line connection to the quantity 

probe inner cylinder, KSC personnel consulted with cognizant personnel 

at MSC and at NR and decided to test whether the oxygen tank no. 2 

could be filled without. problems. 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 between the capacitance plates of the quantity gage would 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 would have taken at least 45 hours. In addition, shelf replacement 

would have had the potential of damaging or degrading other elements of 

the SM in the course of replacement activity. Therefore, the decision 

was made to test the ability to fill oxygen tank no. 2 on March 30, 

1970, twelve 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. 

Accordingly, flow tests with GOX were run on oxygen tank no. 2 

and on oq-gen tank no. 1 for comparison. 

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 low energy levels would result. On the 

filling test, oxygen tanks no. 1 and no. 2 were filled with LOX to 

about 20 percent of capacity on March 30 with no difficulty. 

emptied in the normal manner, but emptying oxygen tank no. 2 again 

required pressure cycling with the heaters turned on. 

No problems were encountered, 

Tank no. 1 

4-22 
As the launch date approached, the oxygen tank no. 2 detanking 

problem was considered by the Apollo organization. At this point, 

the "shelf drop" incident on October 21, 1968, at NR 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. 

portion of the procedure was quite similar, although a slightly lower 

GOX pressure was utilized. 

Throughout these considerations, which involved technical and 

In fact, however, the last 

management personnel of KSC, MSC, NR, Beech, and NASA Headquarters, 

emphasis was directed toward the possibility and consequences of a loose 

fill tube; very little attention was paid to the extended operation of 

heaters and fans except to note that they apparently operated during 

and after the detsnking sequences. 

Many of the principals in the discussions 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 tank, and therefore did not advise management officials 

of any possible consequences of the unusually long heater operations. 

As noted earlier in this chapter, and shown in figure 4-2, each 

heater 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. In tests conducted at MSC since the accident, 

however, it was found that the switches failed to open when the 

heaters were powered from a 65 V dc supply similar to the power used 

at KSC during the detanking sequence. Subsequent investigations have 

sham that the thermostatic switches used, while rated as satisfactory 

for the 28 V dc spacecraft power supply, could not open properly at 

65 V dc. Qualification and test procedures for the heater assemblies 

and switches do not at any time test the capability of the switches 

to open while under full current conditions. A review of the voltage 

recordings made during the detanking at KSC indicates that, in fact, 

the switches did not open when the temperature indication from within 

the tank rose past 80' F. Further tests have shown that the temperatures on the heater tube may have reached as much as 1000° F during 

the detanking. This temperature will cause serious damage to adjacent 

Teflon insulation, and such damage almost certainly occurred. 

None of the above, however, was known at the time and, after 

extensive consideration was given to all possibilities of damage from 

a loose fill tube, it was decided to leave the oxygen shelf and oxygen 

tank no. 2 in the SM and to proceed with preparations for the launch 

of Apollo 13. 

4-23 
The manufacture and test history of oxygen tank no. 2 is discussed 

in more detail in Appendix C to this report. 

. 

4-24 
202 

PART 3. THE APOLLO 13 FLIGHT 

The Apollo l3 mission was designed to perform the third manned 

lunar landing. The selected site was in the hilly uplands of the Fra 

Mauro formation. A package of five scientific experiments was planned 

for emplacement on the lunar surface near the lunar module (LM) landing 

point: (1) a lunar passive seismometer to measure and relay meteoroid 

impact and moonquakes and to serve as the second point in,a seismic net 

begun with the Apollo 12 seismometer; (2) a heat flow device for measuring the heat flux from the lunar interior to the surface and surface 

material conductivity to a depth of 3 meters; (3) a charged-particle 

lunar environment experiment for measuring solar wind proton and electron 

effects on the lunar environment; (4) a cold cathode gage for measuring 

density and temperature variations in the lunar atmosphere; and (5) a 

dust detector experiment. 

Additionally, the Apollo 13 landing crew was to gather the third 

set of selenological samples of the lunar surface for return to earth 

for extensive scientific analysis, 

scheduled to be photographed from lunar orbit with a high-resolution 

topographic camera carried aboard the command module. 

Candidate future landing sites were 

During the week prior to launch, backup Lunar Module Pilot Charles 

M. Duke, Jr., contracted rubella. Blood tests were performed to determine prime crew immunity, since Duke had been in close contact with the 

prime crew. These tests determined that prime Commander James A. Love11 

and prime Lunar Module Pilot Fred Haise were immune to rubella, but that 

prime Command Module Pilot Thomas K. Mattingly 111 did not have immunity. 

Consequently, following 2 days of intensive simulator training at the 

Kennedy Space Center, backup Command Module Pilot John L. Swigert, Jr., was substituted in the prime crew to replace Mattingly. Swigert had 

trained for several months with the backup crew, and this additional 

work in the simulators was aimed toward integrating him into the prime 

crew so that the new combination of crewmen could function as a team 

during the mission. 

Launch was on time at 2:13 p.m., e.s.t., on April 11, 1970, from the 

KSC Launch Complex 39A. The spacecraft was inserted into a 100-nauticalmile circular earth orbit. The only significant launch phase anomaly was 

premature shutdown of the center engine of the S-I1 second stage. As a 

result, the remaining four S-11 engines burned 34 seconds longer than 

planned and the S-IVB third stage burned a few seconds longer than planned. 

of the planned velocity. Moreover, an adequate propellant margin was 

maintained in the S-IVB for the translunar injection burn. 

At orbital insertion, the velocity was within 1.2 feet per second 

4-25 
203 

Orbital insertion was at 00:12:39 ground elapsed time (g.e.t.). 

me initial one and one-half earth orbits before translunar injection 

(TLI) were spent in spacecraft systems checkout and included television 

transmissions as Apollo 13 passed over the Merritt Island Launch Area, 

Florida, tracking station. 

The S-IVB restarted at 02:35:46 g.e.t. for the translunar injection 

burn, with shutdown coming some 5 minutes 51 seconds later. Accuracy of 

the Saturn V instrument unit guidance for the TLI burn was such that a 

planned midcourse correction maneuver at 11:41:23 g.e.t. was not necessary. After TLI, Apollo 13 was calculated to be on a free-return trajectory with a predicted closest approach to the lunar surface of 210 

nautical miles. 

Tne CSM was separated from the S-IVB about 3 hours after launch, 

and after a brief period of stationkeeping, the crew maneuvered the CSM 

to dock with the LM vehicle in the LM adapter atop the S-IVB stage. The 

S-IVB stage was separated from the docked CSM and LM shortly after 4 

hours into the mission. 

In manned lunar missions prior to Apollo 13, the spent S-IVE third 

stages were accelerated into solar orbit by a "slingshot" maneuver in 

which residual liquid owgen was dumped through the J-2 engine to provide propulsive energy. 

stage on the lunar surface in proximity to the seismometer emplaced in 

the Ocean of Storms by the crew of Apollo 12. 

On Apollo 13, the plan was to impact the S-IVB 

Two hours after TLI, the S-IVB attitude thrusters were ground commanded on to adjust the stage's trajectory toward the designated impact 

at latitude 3' S. by longitude 30' W. Actual impact was at latitude 

2.4' S. by longitude 27.9' W.--74 nautical miles from the Apollo 12 

seismometer and well within the desired range. Impact was at 77:56:40 

g.e.t. Seismic signals relayed by the Apollo 12 seismometer as the 

30,700-~ound stage hit the Moon lasted almost 4 hours and provided lunar 

scientists with additional data on the structure of the Moon. 

As in previous lunar missions, the Apollo 13 spacecraft was set up 

in the passive thermal control (PTC) mode which calls for a continuous 

roll rate of three longitudinal axis revolutions each hour. During crew 

rest periods and at other times in translunar and transearth coast when 

a stable attitude is not required, the spacecraft is placed in PTC to 

stabilize the thermal response by spacecraft structures and system. 

At 30:40:49 g.e.t., a midcourse correction maneuver was made using 

the service module propulsion system. The crew preparations for the 

burn and the burn itself were monitored by the Mission Control Center 

(MMC) at MSC by telemetered data and by television from the spacecraft. 

This midcourse correction maneuver was a 23.2 feet per second hybrid 

4-26 
transfer burn which took Apollo 13 off a free-return trajectory and 

placed it on a non-free-return trajectory. 

flown on Apollo 12. 

is to control the arrival time at the Moon to insure the proper lighting 

conditions at the landing site. Apollo 8, 10, and 11 flew a pure freereturn trajectory until lunar orbit insertion. 

transfer maneuver lowered the predicted closest approach, or pericynthion, altitude at the Moon from 210 to 64 nauticd miles. 

A similar trajectory had been 

The objective of leaving a free-return trajectory 

The Apollo 13 hybrid 

From launch through the first 46 hours of the mission, the performAt 46:40:02, the crew turned on the fans in 

ance of oxygen tank no. 2 was normal, so far as telemetered data and 

crew observations indicate. 

oxygen tank no. 2 as a routine operation. Within 3 seconds, the oxygen 

tank no. 2 quantity indication changed from a normal reading of about 

82 percent full to an obviously incorrect reading "off-scale high," of 

over 100 percent. 

shows that this erroneous reading could be caused by either a short circuit or an open circuit in the gage wiring or a short circuit between 

the gage plates. Subsequent events indicated that a short was the more 

likely failure mode. 

Analysis of the electrical wiring of the quantity gage 

At 47:54:50 and at 51:07:44, the oxygen tank no. 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 LM for checkout. At 53:27 g.e.t., the 

Commander (CMR) and Lunar Module Pilot (UP) were cleared to enter the 

LM to commence inflight inspection of the LM. Ground tests before launch 

had indicated the possibility of a high heat-leak rate in the ldil descent 

stage supercritical helium tank. Crew verification of actual pressures 

found the helium pressure to be within normal limits. 

helium is stored in the LM for pressurizing propellant tanks. 

Supercritical 

The LM was powered down and preparations were underway to close the 

LM hatch and run through the presleep checklist when the accident in 

oxygen tank no. 2 occurred. 

At 55:52:30 g.e.t., a master alarm on the CM caution and warning 

system alerted the crew to a low pressure indication in the cryogenic 

hydrogen tank no. 1. This tank had reached the low end of its normal 

operating pressure range several times previously during the flight. 

At 55:52:58, flight controllers in the MCC requested the crew to turn 

on the cryogenic system fans and heaters, 

The command Module Pilot (CMP) acknowledged the fan cycle request 

at 55:53:06 g.e.t., and data indicate that current was applied to the 

oxygen tank no. 2 fan motors at 55:53:20. 

4-27 
About 1-112 minutes later, at 55:54:53.555, telemetry from the 

spacecraft was lost almost totally for 1.8 seconds. During the period 

of data loss, the caution and warning system alerted the crew to a low 

voltage condition on dc main bus B. At about the same time, the crew 

heard a loud "bang" and realized that a problem existed in the 

spacecraft. 

The events between fan turnon at 55:53:20 and the time when the 

problem was evident to the crew and Mission Control are covered in some 

detail in Part 4 of this chapter, "Summary Analysis of the Accident." 

It is now clear that oxygen tank no. 2 or its associated tubing lost 

pressure integrity because of combustion within the tank, and that effects of oxygen escaping from the tank caused the removal of the panel 

covering bay 4 and a relatively slow leak in oxygen tank no. 1 or its 

lines or valves. Photos of the SM taken by the crew later in the mission show 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 inoperative, leaving the CM 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 (used to repressurize the CM after cabin venting), 

cal power and oxygen to permit safe return of the crew to Earth. 

The LM, therefore, became the only source of sufficient electriThe various telemetered parameters of primary interest are shown 

in figure 4-10 and listed in table 4-11. 

4-28 

47-591 0 - 10 - 14 
206 

a 

e 

k 

4-29 
207 

I. 

0 rl 

f 

4-30 
TABU 4-11.- DETAILED CHRONOLOGY FROM 

2.5 MINUTES BEFORE THE ACCIDENT TO 5 MINUTES AFTER THE ACCIDENT 

Time, g.e.t. - Event 

Events During 52 Seconds Prior to First Observed Abnormality 

55: 52: 31 Master caution and warning triggered by low hydrogen 

pressure in tank no. 1. Alarm is turned off after 

4 seconds. 

55: 52:5a Ground requests tank stir. 

55:53:06 Crew acknowledges tank stir. 

55:53:1a Oxygen tank no. 1 fans on. 

55 : 53: 19 Oxygen tank no. 1 pressure decreases 8 psi. 

55: 5320 Oxygen tank no. 2 fans turned. on. 

55 : 53: 20 St abi li zati on control sys tern e le c t ri cal disturbance 

indicates a power transient. 

55: 5321 Oxygen tank no. 2 pressure decreases 4 psi. 

Abnormal Events During 90 Seconds Preceding the Accident 

55: 53:22.718 Stabilization control system electrical disturbance 

indicates a power transient. 

1.2-volt decrease in ac bus 2 voltage. 

11.1-amp rise in fuel cell 3 current for one 

sample. 

Oxygen tank no. 2 pressure begins rise lasting 

for 24 seconds. 

11-volt decrease in ac bus 2 voltage for one 

sample. 

Stabilization control system electrical disturbance 

indicates a power transient. 

55 : 5 3: 22.757 

55: 53: 22.772 

55 : 5 3: 36 

55:53: 38.057 

55:53: 38.085 

4-31 
209 

TABLE 4-11.- DETAILED CHRONOLOGY FROM 

2.5 MINLTES BEFORE THE ACCIDENT TO 5 MINUTES AFTER THE ACCIDENT - Continued 

Time, .q.e.t. 

55 : 53: 41.172 

55:53:41.192 

55 : 54: 31 

55: 54: 43 

55: 54: 45 

55:54:48 

55: 54: 51 

55:54:52 

55 : 54: 52.703 

55 : 54: 53.182 

55:54:53.220 

22.9-amp rise in fuel cell 3 current for one sample. 

Stabilization control system electrical disturbance 

indicates a power transient. 

Oxygen tank no. 2 pressure rise ends at a pressure 

of 953.8 psia. 

Oxygen tank no. 2 pressure begins to rise. 

Oxygen tank no. 2 quantity drops from full scale 

for 2 seconds and then reads 75.3 percent. 

Oxygen tank no. 2 temperature begins to rise 

rapidly . 

Flow rate of oxygen to all three fuel cells begins 

to decrease. 

Oxygen. tank no. 2 pressure reaches maximum value 

of 1008.3 psia. 

Oxygen tank no. 2 temperature rises 40' F for one 

sample (invalid reading). 

Oxygen tank no. 2 quantity jumps to off-scale high 

and then begins to drop until the time of telemetry 

loss, indicating failed sensor. 

Oxygen tank no. 2 temperature reads -151.3' F. 

Oxygen tank no. 2 temperature suddenly goes offscale low, indicating failed sensor. 

Last telemetered pressure from oxygen tank no. 2 

before telemetry loss is 995.7 psia. 

Sudden accelerometer activity on X, Y, and Z axes. 

Stabilization control system body rate changes 

begin. 

4-52 
210 

TABLE 4-11. - DETAILED CHRONOLOGY FROM 

2.5 MINUTES BEF'ORE THE ACCIDENT TO 5 MINUTES AFTER THE ACCIDENT - Continued 

Time, ~.e.t. Event 

55 : 54: 53.323 

55:54:53.5 

55: 54: 53.542 

Oxygen tank no. 1 pressure drops 4.2 psi. 

2.8-amp rise in total fuel cell current. 

X, Y, and Z accelerations in CM indicate l.l7g, 

0.65g and 0.65g, respectively. 

1.8-Second Data Loss 

55:54:53.555 Loss of telemetry begins. 

55:54:53.555+ Master caution and warning triggered by dc main 

bus B undervoltage. 

seconds. All indications are that the cryogenic 

oxygen tank no. 2 lost pressure in this time period 

and the panel separated. 

Nitrogen pressure in fuel cell 1 is off-scale low 

indicating failed sensor. 

Alarm is turned off in 6 

55:54:54.741 

55:54:55.35 Recovery of telemetry data. 

55: 54: 56 

55 : 54: 56 

55:54:56 

55: 54: 56 

Events During 5 Minutes Following the Accident 

Service propulsion system engine valve body temperature begins a rise of 1.65' F in 7 seconds. 

Dc main bus A decreases 0.9 volt to 28.5 volts and 

dc main bus B decreases 0.9 volt to 29.0 volts. 

Total fuel cell current is 15 amps higher than the 

final value before telemetry loss. High current 

continues for 19 seconds. 

Oxygen tank no. 2 temperature reads off-scale high 

after telemetry recovery, probably indicating failed 

sensors. 

Oxygen tank no. 2 pressure reads off-scale low following telemetry recovery, indicating a broken supply 

line, a tank pressure below 19 psi, or a failed sensor. 

4-35 
TPBLE 4-11. - DETAILED CHRONOLOGY FROM 

2.5 MINUTES BEFORE THE ACCIDENT TO 5 MINUTES AFTER THE ACCIDENT - Continued 

55: 54:57 

55: 54:59 

55: 55:Ol 

55 : 55 :02 

55: 55 :02 

55 : 55 : 09 

55: 55:20 

55: 55: 35 

55 : 55 : 49 

55: 56:10 

- Event 

Oxygen tank no. 1 pressure reads 781.9 psia and 

begins to drop steadily. 

Oxygen tank no. 2 quantity reads off-scale high 

following telemetry recovery indicating failed sensor. 

The reaction control system helium tank C temperature 

begins a 1.66" F increase in 36 seconds. 

Oxygen flow rates to fuel cells 1 and 3 approached 

zero after decreasing for 7 seconds. 

The surface temperature of the service module oxidizer tank in bay 3 begins a 3.8' F increase in a 

15-second period. 

The service propulsion system helium tank temperature 

begins a 3.8' F increase in a 32-second period. 

Dc main bus A voltage recovers to 29.0 volts; dc 

main bus B recovers to 28.8 volts. 

Crew reports, "I believe we've had a problem here." 

Crew reports, "We've had a main B bus undervolt." 

Oxygen tank no. 2 temperature begins steady drop 

lasting 59 seconds, probably indicating failed sensor. 

Crew reports, "Okay right now, Houston. The voltage 

is looking good, and we had a pretty large bang 

associated with the caution and warning there. And 

as I recall, main B was the one that had had an amp 

spike on it once before." 

Oxygen tank no. 2 quantity becomes erratic for 69 

seconds before assuming an off-scale-low state, 

indicating failed sensor. 

4-34 
212 

TABLE 4-11. - DETAILED CHRONOLOGY FROM 

2.5 MINUTES BEFORE THE ACCIDENT TO 5 MINUTES AFTER THE ACCIDENT - Concluded 

Time, g.e.t. Event 

55 : 57:04 Crew reports, "That jolt must have rocked the 

sensor on--see now--ovgen quantity 2. 

oscillating down around 20 to 60 percent. 

it's full-scale high again." 

It was 

Nar 

55 : 57: 39 

55:57:40 

55: 57: 44 

55 : 57: 45 

55:57:59 

55:58:02 

55: 58:06 

55:58:07 

55: 58:07 

55 : 56: 25 

56 : 00 : 06 

Master caution and warning triggered by dc main 

bus B undervoltage. 

6 seconds. 

Dc main bus B drops below 26.25 volts and continues 

to fall rapidly. 

Ac bus 2 fails within 2 seconds 

Fuel cell 3 fails. 

Fuel cell 1 current begins to decrease. 

Master caution and warning caused by ac bus 2 

being reset. Alarm is turned off after 2 seconds. 

Master caution and warning triggered by dc main 

bus A undervoltage. Alarm is turned off in 13 

seconds. 

Dc main bus A drops below 26.25 volts and in the 

next few seconds levels off at 25.5 volts. 

Crew reports, 'lac 2 is showing zip." 

Crew reports, "Yes, we got a main bus A undervolt 

now, too, showing. 

Main B is reading zip right now." 

Master caution and warning triggered by high hydrogen 

flow rate to fuel cell 2. Alarm is turned off in 

2 seconds. 

Alarm is turned off in 

It's reading about 25-1/2. 

4-35 
213 

PART 4. SUMMARY ANALYSIS OF THE ACCIDENT 

Combustion in oxygen tank no. 2 led to failure of that tank, damage 

to oxygen tank no. 1 or its lines or valves adjacent to tank no. 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 no. 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 special tests and analyses conducted by the Apollo organization and by or 

for the Board after the accident. (For more information on details of 

mission events, design, manufacture and test of the system, and special 

tests and analyses conducted in this investigation, refer to Appendices 

B, C, D, E, and F of this report.) 

Although tests and analyses are continuing, sufficient information 

is now available to provide a reasonably 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 thus made the wiring susceptible to the electrical short 

circuit which probably initiated combustion within the tank. While the 

exact point of initiation of combustion may never be known with certainty, the nature of the occurrence is sufficiently 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 of combustion, loss of oxygen tank no. 2 system 

integrity, and loss of oxygen tank no. 1 system integrity. 

INITIATION 

Key Data 

55:53:20+ Oxygen tank no. 2 fans turned on. 

55:53:22.757 1.2-volt decrease in ac bus 2 voltage. 

*In evaluating telemetry data, consideration must be given to the 

For further information, reference 

fact that the Apollo pulse code modulation (EM) system samples data in 

time and quantitizes in amplitude. 

may be made to Part B7 of Appendix B. 

4-56 
214 

55:53:22.772 11.1-ampere "spike" recorded in fuel cell 3 current 

followed by drop in current and rise in voltage typical of removal of power from one fan motor--indieating opening of motor circuit. 

55: 53 : 36 Oxygen tank no. 2 pressure begins to rise. 

The evidence points strongly to an electrical short circuit with 

arcing as the initiating event. About 2.7 seconds after the fans were 

turned on in the SM oxygen tanks, an ll..l-ampere current spike and 

simultaneously a voltage-drop spike were recorded in the spacecraft 

electrical system. Immediately thereafter, current dram from the fie1 

cells 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 tank.s at the time and the quantity 

gage and temperature sensor are very lc'w power devices, The next anomalous event recorded was the beginning of a pressure rise in oxygen 

tank no. 2, 13 seconds later. Such a time lag is possible with lowlevel combustion at the time. These facts-point to the likelihood that 

an electrical short circuit with arcinn occurred in the fan motor or its 

leads to initiate the accident sequence. The energy available from the 

short circuit was probably 10 to 20 joules. Tests conducted during 

this investigation have shown that this energy is more than adequate to ignite Teflon of the type contained within the tank. (The 

quantity gage in oxygen tank no. 2 had failed at 46:40 g.e.t. mere 

is no evidence tying the quantity gage failure directly to accident 

initiation, particularly in view of the very low energy available 

from the gage.) 

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 launch. 

Furthermore, there is no evidence :?ointing to any other mechanism 

of initiation. 

PROPAGATION OF COMBUSTION 

Key Data 

55 : 53 : 36 Oxygen tank no. 2 pressure begins rise (same event 

noted previously). 

55: 53:38.057 11-volt decrease recorded in ac bus 2 voltage. 

4-37 
55: 54:OO 

55: 54: 15 

55: 54: 30 

55:54:31 

55 : 54: 45 

55 : 54: 52.763 

22.9-ampere "spike" recorded in fuel cell 3 current, 

followed by drop in current and rise in voltage typical of one fan motor -- indicating opening of another 

motor circuit. 

Oxygen tank no. 2 pressure levels off at 954 psia. 

Oxygen tank no. 2 pressure begins to rise again. 

Oxygen tank no. 2 quantity gage reading drops from 

full scale (to which it had failed at 46:40 g.e.t.) 

to zero and then read 75-percent full. This behavior indicates the gage short circuit may have corrected itself. 

Oxygen tank no. 2 temperature begins to rise rapidly. 

OFjgen tank no. 2 pressure reading reaches maximum 

recorded value of 1008 psia. 

Oxygen tank no. 2 pressure reading had dropped to 

996 psia. 

The available evidence points to a combustion process as the cause 

of the pressure and temperature increases recorded in oxygen tank no. 2. 

The pressure reading for oxygen tank no. 2 began to increase about l3 

seconds after the first electrical spike, and about 55 seconds later the 

temperature began t3 increase. The terqerature sensor reads local temperature, which need not represent bulk fluid temperature. Since the 

rate of pressure rise in the tapk indicates a relatively slow propagation of burning, it is likely that the region immediately around the 

temperature sensor did not become heated until this time. 

There are materials within the tank that can, if ignited in the 

presence of supercritical oxygen, react chemically with the oxygen in 

exothermic chemical reactions. The most readily reactive is Teflon 

used for electrical insulation in the tank. Also potentially reactive 

are metals, particularly aluminum. 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 would be more likely characteristic of Teflon combustion 

than metal-oxygen reactions. 

While the data available on the combustion of Teflon in supercritical oxygen in zero-g are extremely limited, those which are available 

indicate that the rate of combustion is generally consistent witin these 

4-38 
216 

observations. The cause of the 15-second period of relatively constant 

pressure fbst indicated at 55:53:59.763 has not been precisely determined; it is believed to be associated with a change in reaction rate as 

combustion proceeded through various Teflon elements. 

While there is enough electrical power in the tank to cause ignition 

in the event of a short circuit or abnormal heating in defective wire, 

there is not sufficient electric power to account for all of the energy 

required to produce the observed pressure rise. 

LOSS OF OXYGEN TANK NO. 2 SYSTEM IKTEGRITY 

Key Data 

55: 54: 52 Last valid temperature indication (-151" F) from 

oxygen tank no. 2. 

Last pressure reading from oxygen tank no. 2 before 

loss of data--996 psia. 

Sudden accelerometer activity on X, Y, and Z axes. 

Stabilization control system body rate changes begin. 

Loss of telemetry data begins. 

Recovery of telemetry data. 

Various temperature indications in SM begin slight 

rises. 

Oxygen tank no. 2 temperature reads off-scale high. 

Oxygen tank no. 2 pressure reads off-scale low. 

55 : 54: 52.763 

55: 54: 53.182 

55 : 54: 53.220 

55: 54: 53.555* 

55:54: 55.35 

55:54:56 

55: 54: 56 

55:54:56 

After the relatively slow propagation process described above took 

place, there was a relatively abrupt loss of oxygen tank no. 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, 

for 8 seconds, dropping to 996 psia before readings were lost. 

Pressure began a decrease 

Virtually 

*Several bits of data have been obtained from this "loss of telemetry data" period. 

4-39 
217 

all signals from the spacecraft were lost about 1.85 seconds after the 

last presumably valid reading from within the tank, a temperature reading, and 0.8 second after the last presumably valid pressure 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). 

Abnormal spacecraft accelerations were recorded approximately 0.42 second after the last pressure reading and approximately 0.38 second before 

the loss 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 probably 

occurred in this time period. Telemetry signals from Apollo 13 were 

lost for a period of 1.8 seconds. 

ment indicators from oxygen tank no. 2 were off-scale, high or low. Temperatures recorded by sensors in several different locations in the SM 

showed slight increases in the several seconds following reacquisition 

of signal. 

jettisoned show that the bay 4 panel was ejected, undoubtedly during 

this event. 

When signal was reacquired, all instruPhotographs taken later by the Apollo 13 crew as the 34 was 

Data are not adequate to determine precisely the way in which the 

oxygen tank no. 2 system lost its integrity. However, available information, analyses, and tests performed during this investigation indicate 

that most probably the combustion within the pressure vessel ultimately 

led to localized heating and failure at the pressure vessel closure. It 

is at this point, the upper end of the quantity probe, that the l/2-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 location where all of the 

wires come together. This, possibly augmented by ignition of the 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 lead immediately to pressurization of 

the tank dome, which is equipped with a rupture disc rated at about 75 

psi. Rupture of this disc or of the entire dome would then release 

oxygen, accompanied by combustion products, into bay 4. The accelerations recorded were probably caused by this release. 

Release of the oxygen then began to pressurize the oxygen shelf 

space of bay 4. 

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 was accompanied by combustion of Mylar and 

Kapton (used extensively as thermal insulation in the oxygen shelf compartment, figure 4-11, and in the tank done) which would augment the 

If the hole formed in the pressure vessel were large 

4-40 
218 

4-11. - Figure Xh-YXX Closeup view of oxygen tank shelf. 

4-41/4-42 
219 

pressure caused by the oxygen itself. The slight temperature increases 

recorded at various SM locations indicate that combustion external to 

the tank probably took place. Further testing may shed additional light 

on the exact mechanism of panel ejection. The ejected panel then struck 

the high-gain antenna, disrupting communications from the spacecraft for 

the 1.8 seconds. 

LOSS OF OXYGEN Tf!JK NO. 1 INTEGRITY 

Key Data 

55 :54:53.323 Oxygen tank no. 1 pressure drops 4 psia (from 883 psia 

to 879 psia). 

55:54:53.555 LOSS of telemetry data. to 

55 :54:55.35 

55: 54~56 Oxjrgen tank no. 1 pressure reads 782 psia and drops 

steadily. Pressure drops over a period of l3O minutes to the point at which it was insufficient to 

sustain operation of fuel cell no. 2. 

There is no clear evidence of abnormal behavior associated with 

oxygen tank no. 1 prior to loss of signal, although the one data bit 

(4 psi) drop in pressure in the last tank no. 1 pressure reading prior 

to loss of signal may indicate that a problem was beginning. Immediately 

after signal strength was regained, data show that tank no. 1 system had 

lost its integrity. 

approximately l3O minutes, indicating that a relatively slow leak had 

developed in the tank no. 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 or relief valve. 

Pressure decreases were recorded over a period of 

Since there is no evidence that there was any anomalous condition 

arising within oxygen tank no. 1, it is presumed that the loss of oxygen 

tank no. 1 integrity resulted from the oxygen tank no. 2 system failure. 

The relatively sudden, and possibly violent, event associated with loss 

of integrity of the oxygen tank no. 2 system could have ruptured a line 

to oxygen tank no. 1, or have caused a valve to leak because of mechanical shock. 

4-43 
PART 5. APOLLO 13 RECOVERY 

UNDERSTANDING THE PROBLEM 

In the period immediately following the caution and warning alarm 

for main bus 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 loss of telemetered data was accompanied by the 

switching of the CSM high-gain antenna mounted on the SM adjacent to 

bay 4 from narrow beam width to wide beam width. 

does this automatically 200 milliseconds after its directional lock on 

the ground signal has been lost. 

The high-gain antenna 

A confusing factor was the repeated firings of various SM attitude 

control thrusters during the period after data loss. 

these thrusters were being fired to overcome the effects that oxygen 

venting and panel blowoff were having on spacecraft attitude, but it 

was believed for a time that perhaps the thrusters were malfunctioning. 

The failure of oxygen tank no. 2 and consequent removal of the bay 4 

In all probability, 

panel produced a shock which closed valves in the oxygen supply lines to 

fuel cells 1 and 3. These fuel cells ceased to provide power in about 3 

minutes, when the supply of oxygen between the closed valves and the 

cells was depleted, Fuel cell 2 continued to parer ac bus 1 through dc 

main bus A, but the failure of fuel cell 3 left dc main bus B and ac 

bus 2 unparered (see fig. 4-12). The oxygen tank no. 2 temperature and 

quantity gages were connected to ac bus 2 at the time of the accident. 

Thus, these parameters could not be read once fuel cell 3 failed at 

55:57:44 until power was applied to ac bus 2 from main bus A. 

The crew was not alerted to closure of the oxygen feed valves to 

fuel cells 1 and 3 because the valve position indicators in the CM were 

arranged to give warning only if both the oxygen and hydrogen valves 

closed. The hydrogen valves remained open. The crew had not been 

alerted to the oqygen tank no. 2 pressure rise or to its subsequent drop 

because a hydrogen tank low pressure warning had blocked the cryogenic 

subsystem portion of the caution and warning system several minutes before the accident. 

When the crew heard the bang and got the master alarm for low dc 

main bus B voltage, the Commander was in the lower equipment bay of the 

command module, stowing a television camera which had just been in use. 

4-44 
f 

c 

Ul 

I Inverter 

1 DC mainA -

tank 2 

pressure 

Figure 4-12.- Electrical 

configurat
222 

The Lunar Module Pilot was in the tunnel between the CSM and the LM, 

returning to the CSM. The Command Module Pilot was in the left-hand 

couch, monitoring spacecraft performance. Because of the master alarm 

indicating low voltage, the CMP moved across to the right-hand couch 

where CSM voltages can be observed. He reported that voltages were 

"looking good" at 55:56:10. 

fuel cell 3 did not fail for another 1-1/2 minutes. 

fluctuations in the oxygen tank no. 2 quantity, followed by a return 

to the off-scale high position. 

ment) . 

At this time, main bus B had recovered and 

He also reported 

(See fig. h-13 for CM panel arrangeWhen fuel cells 1 and 3 electrical output readings went to zero, 

the ground controllers could not be certain that the cells had not somehow been disconnected from their respective busses and were not othemise 

all right. Attention continued to be focused on electrical problems, 

Five minutes af'ter 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 was realized that fuel cells 1 and 3 were 

not functioning, the crew was directed to perform an emergency powerdown 

to lower the load on the remaining fuel cell. Observing the rapid decay 

in oxygen tank no. 1 pressure, controllers asked the crew to switch power 

to the oxygen tank no. 2 instrumentation. When this was done, and it 

was realized that oxygen tank no. 2 had failed, the extreme seriousness 

of the situation became clear. 

During the succeeding period, efforts were made to save the remaining oxygen in the oxygen tank no. 1. Several attempts were made, but 

had no effect. The pressure continued to decrease. 

It was obvious by about 1-1/2 hours after the accident that the 

oxygen tank no, 1 leak could not be stopped Bnd that shortly it would be 

necessary to use the LM as a "lifeboat" for the remainder of the mission. 

By 58:40 g.e.t., the LM had been activated, the inertial guidance 

reference transferred from the CSM guidance system to the LM guidance 

system, and the CSM systems were turned off. 

RETURN TO EARTH 

The remainder of the mission was characterized by two main activities--planning and conducting the necessary propulsion maneuvers 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 actual control vehicle for the time required. 

4-46 

224 
225 

One significant anomaly was noted during the remainder of the 

mission. At about 97 hours 14 minutes into the mission, the LMP 

reported hearing a "thump" and observing venting from the LM. 

data review shows that the LM electrical power system experienced a 

brief but major abnormal current flow at that time. 

that this anomaly was related to the accident. Analysis by the Apollo 

organization is continuing. 

Subsequent 

There is no evidence 

A number of propulsion options were developed and considered. It 

was necessary to return the spacecraft to a free-return trajectory and 

to make any required midcourse corrections. Normally, the service propulsion system (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 SM after the accident, it was decided to use 

the LM descent engine if possible. 

The minimum practical return time was 133 hours g.e.t. to the 

Atlantic Ocean, and the maximum was 152 hours g.e.t. to the Indian 

Ocean. Recovery forces were deployed in the Pacific. The return path 

selected was for splashdown in the Pacific Ocean at 142:40 g.e.t. This 

required a minimum of two burns of the LM descent engine. A third burn 

was subsequently made to correct the n6rmal 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 

4-14 and 4-15 depict the flight plan followed from the time of the accident to splashdown. 

The most critical consumables were water, used to cool the CSM and 

LM systems during use; CSM and LM battery power, the CSM batteries being 

for use during reentry and the LM batteries being needed f o r t h e rest 

of the mission; LM ovgen for breathing; and lithium hydroxide (LiOH) 

filter cannisters used to remove carbon dioxide from the spacecraft 

cabin atmosphere. These consmables, and in particular the water and 

LiOH cannisters, appeared to be extremely marginal in quantity shortly 

after the accident, but once the,LM was powered dam to conserve electric 

power and to generate less heat and thus use less water, the situation 

improved greatly. 

crew to use materials on board to fashion a device allcwing use of the 

CM LiOH cannisters in the LM cabin atmosphere cleaning system (see 

fig. 4-16). At splashdown, many hours of each consumable remained 

available (see figs. 4-17 through 4-19 and table 4-111). 

Engineers at MSC developed a method which allowed the 

4-48 
\ 

- /--- 

1( 

f 

-I=- 

u3 

for e

corri

(105

for entry 

corridor 

( 137 :4 0) 

Figure 4-14.- Translunar trajecto
- 

r CM Dower UP 

G. E:T. 140110 

\/ LM jettison 

G.E.-T. 141:30 

D 

L Landinq 

G. E.T. 14~54 

Figure 4-15.- Final trajector
228 

Figure 4-16.- Lithium hydroxide canister modification. 

4-51 
229 

0 

0 

m 

0 

In 

N 

0 

In 

0 

In rl 

0 

f 

0 

2 

0 

N 

rl 

0 

rl 

rl 

0 

0 

rl 

0 

0. 

0 

m 

0 

b 

0 

9 

4- 52 
230 

I I I I I I I I I 

0 8 0 0 0 0 0 0 0 0 

N d d 0 9 N co P 

(du!u!eurar alqesn) sry duy 

4-53 
23 1 

0 0 0 0 0 0 

9 In d m N rl 

sq1 'ua6Axo 

0 

b 

0 

In 

rl 

0 

d 

d 

0 

W 

rl 

0 

N 

rl 

3 

4 

d 

3 

2 

4 

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> 

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3 

. 

3 

0 

3 

7 

a7 aJ 

% 

s I- 

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! 

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k 

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2l 

4-54 
232 

TABLE 4-111.- CABIN ATMOSPHERE CARBON DIOXIDE 

REMOVAL BY LITHIUM HYDROXIDE 

Required 

Available in LM 

Available in CM 

85 hours 

53 hours 

182 hours 

A more detailed recounting of the events during the Apollo 13 

launch countdm and mission will be found in Appendix B to this report. 

4-55 
233 

CHAPTER 5 

FINDINGS, DETERMINATIONS, AND RECOMICENDATIONS 
PART 1. INTRODUCTION 

The following findings, determinations, and recommendations are the 

product of about 7 weeks of concentrated review of the Apcllo 13 accident 

by the Apollo 13 Review Board. 

accident investigation by the Manned Spacecraft Center (MSC) and its contractors, and on an extensive series of special tests and analyses performed by or for the Board and its Panels. 

They are based on that review, on the 

Sufficient work has been done to identify and understand the nature 

of the malfunction and the direction which the corrective actions must 

take. All indications are that an e l e c t r i c a l l y i n i t i a t e d fire in oxygen 

tank no. 2 in the service module (SM) was the cause of the accident. Accordingly, the aoard has concentrated on this tank; on its design, manufacture, test, handling, checkout, use, failure mode, and eventual effects 

on the rest of the spacecraft. 

the most pr bable cause has been identified. However, at the time of this 

report, sop details of the accident are not completely clear. 

fie accident is generally understood, and 

Fwther tests and analyses, which will be carried out under the overall direction of MSC, will continue to generate new information relative 

to this accident. It is possible that this evidence may lead to ccnclusions differing in detail from those which can be drawn now. However, it 

is most unlikely that fmdamentally different results will be obtained. 

Recommendations are provided as to the general direction which the 

corrective actions should take. Significant modifications should be made 

to the SM 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 

necessary for the Board to review critically the equipment and the organi- 

.zational 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 unusual 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 no. 2 

which flew on Apollo 13 was shipped from Beech Aircraft Corporation to 

North American Rockwell (NR) in apparently satisfactory condition. 

b. It is now known, however, that the tank contained two protective 

thermostatic switches on the heater assembly, wnich were inadequate and 

would subsequently fail during ground test operations at Kennedy Space 

Center (KSC). 

5-1 
c. In addition, it is probable thar. the tank contained a loosely 

fitting fill tube assembly. 

subsequent handling, which included an incident at the prime contractor's 

plant in which the tank was jarred. 

,This assembly was probably displaced during 

d. In itself, the displaced fill tube assembly was not particularly 

serious, but it led to the use of improvised detanking procedures 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 no. 2 using 

normal procedures at KSC. Tests and analyses indicate that this was due 

to gas leakage through tine displaced fill tube assembly. 

f. The special detanking procedures at KSC subjected the tank to an 

extended period of heater operation and pressure cycling. These procedures had not been used before, and the tank had not been qualified by 

test for the conditions experienced. However, the procedures did not 

violate the specifications which governed the operation of the heaters at 

KSC , 

g. In reviewing these procedures before the flight, officials of 

KASA, 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 tc protect the tank. 

h. A number of factors contributed to the presence of inadequate 

thermostatic switches in the heater assembly. The original 1962 specifications from NR to Beech Aircraft Corporation for the tank and heater 

assembly specified the use of 28 V dc power, which is used in the spacecraft. In 1965, NR 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 ar; XSC to reduce pressurization time. Beech 

orjered switckes 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 NASA, N2, 

or Beech In their review of dcc.mentation, nor dld tests identify the incorpatibility of the switches with the gro-und support equipment (GSE) at 

KSC. sir.ce 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, 

2. ?"re thermostatic switches could accommodate the 65 V dc during 

Lank pressurization because they normally remained cool and closed. However, they could not open without damage with 65 V dc power applied. 

were ?.e7er reqJire3 to do so until the special detanking. 

They 

During this 

5-2 
procedure, as the switches started to open when they reached their upper 

temperature limit, they were welded permanently closed by the resulting 

arc and were rendered inoperative as protective thermostats. 

k. Failure of the thermostatic switches to open could have been 

detected at KSC if switch operation had been checked by observing heater 

current readings on the oxygen tank heater control panel. 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. 

Although it 

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 8-hour period of heater 

operation. Such heating has been shown by tests to severely damage the 

Teflon insulation on the fan motor wires 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 wiring, possibly moved by the fan stirring, short circuited 

and ignited its 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, 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 13, the Board has also 

attempted to identify those additional technical and management 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, 

in equipment and procedures that the Board has identified. 

nature of a review board report. 

The report highlights in detail faults or deficiencies 

This is the 

5-3 
237 

It is important, however, to view the criticisms in 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. 

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. 

It has flown to the Moon five times and landed 

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. 

5-4 

47-591 0 - 70 - 16 
238 

1. Findings 

a. 

b. 

c. 

d. 

e. 

f. 

g. 

h. 

i. 

PART 2. ASSESSMENT OF ACCIDENT 

FAILURE OF OXYGEN ThYK NO. 2 

The Apollo 13 mission was aborted as the direct result of 

the rapid loss of oxygen from oxygen tank no. 2 in the SM, 

followed by a gradual loss of oxygen from tank no. 1, and 

a resulting loss of power from the oxygen-fed fuel cells, 

There is no evidence of any forces external to oxygen tank 

no. 2 during the flight which mighr; have cause9 its failire. 

Oxygen tank no. 2 contained materials, including Teflon and 

alminum, which if ignited will burn in supercritical 

oxygen. 

Oxygen tank no. 2 contained potential ignition sources: 

electrical wiring, unsealed electric motors, and rotating 

alminum fans. 

During the.specia1 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 V dc in order to protect the heaters from overneating. The switches were only rated at 30 V dc and have 

been shown to weld closed at the higher volsage. 

Data indicate that in flight the tank heaters located in 

oxygen tanks no. 1 and no. 2 operated normally prior to the 

accident, and they were not on at the time of the accident. 

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 millijoules per 

second. 

Telemetry data immediately prior to the accident indicate 

electrical disturbances of a character which would be caused 

by short circuits accompanied by electrical arcs in the fan 

motor or its leads in oxygen tank no. 2. 

The pressure and temperature within oxygen tank no. 2 rose 

abnormally during the 1-1/2 minutes immediately prior to the 

accident. 

5- 5 
239 

Determinations 

The cause of the failure of oxygen tank no. 2 was combustion 

within the tank. 

Analysis showed that the electrical energy flowing into the 

tank could not account for the observed increases in pressure 

and temperature. 

The heater, temperature sensor, and quantity probe did not 

initiate the accident sequence. 

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. 

The protective thermostatic switches on the heaters in 

oxygen tank no. 2 failed closed during the initial portion 

of the first special detanking operation. This subjected 

the wiring in the vicinity of the heaters to very high temperatures which have been subsequently shown to severely 

degrade Teflon insulation. 

The telemetered data indicated electrical arcs of sufficient 

energy t o ' i g n i t e the Teflon insulation, as verified by subsequent tests. Tnese tests also verified that the 1-ampere 

fuses on the fan motors would pass sufficient energy to ignite the insulation by the mechanism of an electric arc. 

The combustion of Teflon wire insulation alone could release 

sufficient heat to account for the observed increases 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 derconstrated by subsequent tests. 

Tne rate of flame propagation along Teflon-insulated wires 

as measured in subsequent tests is consistent with the indicated rates of pressure rise within the tank. 

SECOivTARY EFFECTS OF TANK FAILURE 

2. Findings 

a. Failure of the tank was accompanied by several events including: 

5-6 
A 'bang" as heard by the crew. 

Spacecraft motion as felt by the crew and as measured by 

the attitude control system and the accelerometers in the 

command module (CM) I 

Momentary loss of telemetry. 

Closing of several valves by shock loading. 

Loss of integrity of the oxygen tank no. 1 system. 

Slight temperature increases in bay 4 and adjacent sectors 

of the SM. 

Loss of the panel covering bay 4 of the SM, as observed and 

photographed by the crew. 

Displacement of the fuel cells as photographed by the crew. 

Damage to the high-gain antenna as photographed by the crew. 

The panel covering of bay 4 could be blown off by pressurization of the bay, 

is required to blow off the panel. 

b. 

About 25 psi of uniform pressure in bay 4 

c. me various bays and sectors of the SM are interconnected 

wit'n open passages so that all would be pressurized if any 

one were supplied with a pressurant at a relatively slow 

rate, 

d. The CM attachments would be failed by an average pressure of 

about 10 psi on the CM heat shield and this would separate 

the CM from the SM. 

Determinations 

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

pressurization of bay 4 of the SM by the high-pressure 

oxygen that escaped from the tank. This pressure pulse may 

have blown off the panel covering bay 4. was substantiated by a series of special tests. 

The pressure pulse from a tank failure might have been 

augmented by combustion of Mylar or Kapton insulation or 

both when subjected to a stream of ozfgen and hot particles 

emerging from the top of the tank, as demonstrated in subsequent tests. 

This possibility 

(2) 

5-7 
(3) Combustion or vaporization of the Mylar or Kapton might 

account for the discoloration of the SM engine nozzle as 

observed and photographed by the crew. 

Photographs of the SM by the crew did not establish the 

condition of the oxygen tank no. 2. 

The high-gain antenna damage probably resulted from striking 

by the panel, or a portion thereof, as it left the SM. 

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

loss of power resulted from the tank no. 2 failure. 

Telemetry, although good, is insufficient to pin down the 

exact nature, sequence, and location of each event of the 

accident in detail. 

(4) 

(5) 

(6) 

(7) 

(8) The telemetry data, crew testimony, photographs, and special 

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

OXYGEN TAM( NO. 2 DESIGN 

3. Findings 

a. The cryogenic oxygen storage tanks contained a combination 

of oxidizer, combustible material, and potential ignition 

sources. 

b. Supercritical oxygen was used to minimize the weight, 

volume, and fluid-handling problems of the oxygen supply 

sys tem. 

c. The heaters, fans, and tank instrumentation are used-in the 

measurement and management of the oxygen supply. 

Determinations 

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

Apollo system. 

Heaters are required to maintain tank pressure as the oxygen 

supply is used. 

Fans were used to prevent excessive pressure drops due to 

stratification, to mix the oxygen to improve accuracy of 

(2) 

(3) 

5-8 
242 

quantity measurements, and to insure adequate heater input 

at,low densities and high oxygen utilization rates. 

need for oxygen stirring on future flights requires father 

investigation. 

The amount of material in the tank which could be ignited 

and burned in the given environment could have been reduced 

signiflcanzly. 

The potential ignition sources constituted an undue hazard 

when considered in the light of the particular tank design 

with its assenbly difficulties. 

YASA, the prime contractor, and the supplier of t2e tank 

were noc fiully aware of the extent of this hazard. 

Zxamlnation of the high-pressure oxygen system in the service 

nod-de following the Apollo 204 fire, which directed attention to the danger of fire in a pure oxygen enviroment, 

failed 20 recognize the deficiencies of the Lank. 

The 

(4) 

(5) 

(6) 

(7) 

P,Tr'LIGBi DAMAGE TO TMX WIRIKG 

4. Findings 

a. Tne oxygen tank no. 2 heater assembly contained two thermostatic switches designed to protect the heaters from overheazing. 

c. 3e thermostatic switches were designed to open and interrupt 

the heater urrent at 80" C 10" P. 

The heaters are operated on 28 V dc in flight and at NX. 

The heaters are operated on 65 V ac at Beech Aircraft Corporation and 65 V dc at the Kennedy Space Center. 

higner voltages are used to accelerate tank pressurization, 

c. 

d. 

Yqese 

e. Yfle Lhermostatic switches were rated at 7 amps at 30 'J dc. 

Mile 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. 

f. Xeither qualification nor acceptance testing of the heater 

The only test of switch opening 

assemblies or the tanks required thermostatic switch opening 

to be checked at 65 V dc. 

5-9 
g. 

h. 

i. 

j. 

k. 

1. 

m. 

was a continuity check at Beech in which the switch was 

cycled open and closed in an oven. 

The thermostatic switches had never operated in flight because this would only happen if the oxygen supply in a tank 

were depleted to nearly zero. 

The thermostatic switches had never operated on the ground 

under load because the heaters had only been used with a 

relatively full tank which kept the siiitches cool and closed. 

During the CDDT, the oxygen tank no. 2 would not detank in 

a normal manner, 

procedure was followed which subjected the heater to about 

8 hours of continuous operation until the tanks were nearly 

depleted of oxygen. 

A second special detanking of shorter duration followed on 

March 30, 1970. 

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. 

The recorded internal tank temperature went off-scale high 

early in the special detanking. The thermostatic switches 

would normally open at this point but the electrical records 

show no thermostatic switch operation. These indications 

were not detected at the time. 

The oxygen tank heater controls at KSC contained ammeters 

which would have indicated thermostatic switch operation. 

On March 27 and 28, a special detanking 

Determinations 

During the special detanking of March 27 and 28 at KSC, when 

the heaters in oxygen tank no. 2 were left on for an extended 

period, the thermostatic switches started to open while 

po-wered by 65 V dc and were probably welded shut. 

Failure of the thermostatic switches to open could have been 

detected at KSC if switch operation had been checked by 

observing heater current readings on the oxygen tank heater 

control panel. Although it was not recognized at the time, 

the tank temperature readings indicated that tke heaters ha6 

reached their temperature limit and switch opening should 

have been expez’ed. 

5-10 
(3) The fact that the switches were not rated to open at 65 V dc 

was not detected by NASA, Mi, or Beech in their reviews of 

documentation or in qualification and acceptance testing. 

(4) The failed switches resulted in severe overheating. Subsequent tests showed that heater assembly temperatures could 

have reached about 1000° F. 

(5) The high temperatures severely damaged the Teflon insulation on the wiring in the vicinity of the heater assembly and set 

the stage for subsequent short circuiting. As shown in 

subsequent tests, this damage could range from cracking to 

total oxidation and disappearance of the insulation. 

During and following the special detanking, the oxygen tank 

no. 2 was in a hazardous condition whenever it contained 

oxygen and was electrically energized. 

(6) 

5-11 
245 

PART 3. SUPPORTING CONSIDERATIONS 

DESIGN, MANUFACTURING, AND TEST 

3. Finding 

The pressure vessel of the supercritical oxygen tank is con- structed of Inconel 718, and is moderately stressed at normal 

operating pressure. 

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 immersed in supercritical oxygen. 

b. Fan motors of this design have a test history of failure 

during acceptance test which includes phase-to-phase and 

phase-to-ground faults. 

c. The fan motor stator windings are constructed with Tefloncoated, ceramic-insulated, number 36 AWG wire. Full phaseto-phase and phase-to-ground insulation is not used in the 

motor design. 

d. The motor case is largely aluminum. 

Determinations 

(1) The stator winding insulation is brittle and easily fractured 

during manufacture of the stator coils. 

The use of these motors in supercritical oxygen was a questionable practice. 

(2) 

7. Findings 

a, The cryogenic oxygen storage tanks contained materials that 

could be ignited and which will burn under the conditions 

5-12 
b. 

C. 

d. 

e. 

f. 

prevailing within the tank, including Teflon, aluminum, 

solder, and Drilube 822. 

The tank contained electrical wiring exposed to the supercritical oxygen. 

Some wiring was in close proximity to heater elements and 

to the rotating fan. 

The design was such that the assembly of the equipment was 

essentially "blind" and not amenable to inspection after 

completion. 

Teflon insulation of the electrical wiring inside the cryogenic oxygen storage tanks of the SM was exposed to relatively sharp metal edges of tank inner parts during manufacturing assembly operations. 

Portions of this wiring remained unsupparted in the tank on 

conpletion of assembly. 

The wiring was insulated with Teflon. 

Determinations 

Tne tank contained a hazardous combination of materials and 

potential ignition sources. 

Scraping of the electrical wiring insulation against metal 

inner parts of the tank constituted a substantial cumulative 

hazard dilring assembly, handling, test, checkout, and operational use. 

"Cold flow" of the Teflon insulation, when pressed against 

metal corners within the tank for an extended period of 

time, could result in an eventual degradation of insulation 

protection. 

The externally applied electrical tests (500-volt Hi-pot) 

could not reveal the extent of such possible insulation 

damage but could only indicate that the relative positions 

of the wires at the time of the tests were such that the 

separation or insulation would withstand the 500-volt potential without electrical breakdown. 

The design was such that it was difficult to insure against 

these hazards. 

There is no evidence that the wiring was damaged during manufacturing. 

5-13 
247 

9. Findings 

a. Dimensioning of the short Teflon and Inconel tube segments 

of the cryogenic oxygen storage tank fill line was such that 

looseness to the point of incomplete connection was 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. 

Probing with a hand tool was used in manufacturing to compensate for limited visibility of the tube segment positions, 

c. 

Determination 

It was possible for a tank to have been assembled with a set Gf 

relatively loose fill tube parts that could go undetected in 

final inspection and be subsequently displaced. 

10. Findings 

a. me Apollo spacecraft system contains numerous pressure 

vessels, many of which carry oxidants, plus related valves 

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 Manned Spacecraft Center. 

c. he piece of equipment, the fuel cell oxygen supply valve 

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. 

DeterminatiGn 

The fuel cell oxygen supply valve module has been identified as 

potentially hazardous. 

11. Findings 

a. in the normal sequence of cryogenic oxygen storage tank integration and 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 34. 

5-14 
248 

b. The SM undergoes factory transportation, air shipment to KSC, 

and subsequent ground transportation and handling. 

Determination 

There were 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 remove the oxygen shelf assembly from SM 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 valves. 

The orygen shelf assembly incorporating S/N XTA0008 in the 

tank no, 2 position, which had been shaken during removal 

from SM 106, was installed in SM 109 one month later. 

c. 

d. An analysis, shelf inspection, and a partial retest emphasizing electrical continuity of internal wiring were accomplished before reinstallation. 

Determinations 

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

the "shelf drop" incident at the prime contractor's plant, 

without detection. 

Other damage to the tank may have occurred from the jolt, 

but special tests and analyses indicate that this is unlikely. 

(2) 

(3) The "shelf drop" incident was not brought to the attention 

of project officials during subsequent detanking difficulties 

at KSC. 

13. Finding 

Detanking, expulsion of liquid oxygen out the fill line of the 

oxygen tank by warm gas pressure applied through the vent line, 

was a regular activity at Beech Aircraft, Boulder, Colorado, in 

emptying a portion of the oxygen used in end-item acceptance 

tests. 

5-15 
249 

Determination 

The latter stages of the detanking operation on oxygen tank 

no, 2 conducted at Beech on February 3, 1967, were similar to 

the standard procedure followed at KSC during the CDDT. 

14. Findings 

a, The attempt to detank 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. 

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. 

It was employed both after CDDT and after a special test to 

verify that the tank could be filled. 

b. 

c. 

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 completion of detanking 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 showed that the current 

associated with the KSC 65 V dc ground powering of the 

heaters would cause the thermostatic switch contacts to 

weld closed if they attempted to interrupt this current. 

g. A 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 wires. Tnese temperatures 

are high enough to damage Teflon and melt solder, 

Determinations 

(1) Oxygen tank no. 2 (XTA 0008) did not detank after CDDT in a manner comparable to its performance the last time it had 

contained liquid oxygen, i.e., in acceptance test at Beech. 

Such evidence indicates that the tank had undergone some 

change of internal configuration during the intervening 

events of the previous 3 years. 

(2) 

5-16 
250 

The tank conditions during the special detanking procedures 

were outside all prior testing of Apollo CSM cryogenic oxygen 

storage tanks. Heater assembly temperatures measured in subsequent tests exceeded 1000" F. 

Severe damage to the insulation of electrical wiring internal 

to %he tank, as determined from subsequent tests, resulted 

from the special procedure. 

Damage to the insulation, particularly on the long un- supported lengths of wiring, may also have occurred due to 

boiling associated with this procedure. 

MSC, KSC, and ITR 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 Apollo 13. 

b. Launch operations personnel who made the change did not have 

a detailed understanding of the tank internal components, or 

the tank history. Tney made appropriate contacts before 

making %he change. 

c. Communications, primarily by telephone, among MSC, KSC, NR, 

and Beech personnel during final launch preparations regarding the cryogenic oxygen system included incomplete and 

inaccurate information. 

d. The MSC 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 65 

to 85 V dc and 9 to 17 amperes with no restrictions on time. 

Determinations 

(1) NR and MSC personnel who prepared %he TSCD did not know that 

the tank heater thermostatic switches would not protect 

the tank. 

Launch operations personnel assumed the tank was protected 

from overheating by the switches. 

(2) 

5-17 
(3) Launch operations personnel at KSC stayed within the 

specified tank heater voltage and current limits during the 

detanking at KSC. 

16. Findings 

a. After receipt of the Block I1 oxygen tank specifications 

from KR, which required the tank heater assembly to operate 

with 65 V dc GSE power only during tank pressurization, Beech 

Aircraft did not require their Block I thermostatic switch 

supplier to make a change in the switch to operate at the 

higher voltage. 

b. NR did not review the tank or heater to assure compatibility 

between the switch and the GSE. 

c. MSC did not review the tank or heater to assure compatibility between the switch and the GSE. 

d. No tests were specified by MSC, NR, or Beech to check this 

switch under load. 

Determinations 

(1) NR and Beech specifications governing the powering and the 

thermostatic switch protection of the heater assemblies were 

inadequate. 

The specifications governing the testing of the heater 

assemblies were inadequate. 

(2) 

17. Findirg 

The hazard associated with the long heater cycle during detanking 

was not given consideration in the decision t o f l y oxygen tank 

no. 2. 

Determinations 

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

thermostatic switches did not protect the tank from overheating, 

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

thermostatic switches had been known, the tank would have 

been replaced. 

5-18 
252 

18. Findings 

a. 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. 

When the Apollo 13 cryogenic oxygen system was originally 

designed, the management controls were not defined in as 

great detail as they are now. 

b. 

Determination 

From reyiew 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 

Apollo 13. 

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

was a small leak through the vent quick disconnect, which was 

corrected. 

Determination 

No indications of a potential inflight malfunction of the oxygen 

tank no. 2 were present during the launch countdown. 

MISSION EVENTS THROUGH ACCIDENT 

20. Findings 

a. 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 CM. 

Determinations 

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

of the Board except insofar as it relates to the Apollo 13 

accident. 

5-19 
253 

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

probably did not affect the oxygen tank. 

21. Findings 

a, Fuel cell current increased between 46:40:05 and 46:40:08 

indicating that 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. 

Determinations 

(1) The oxygen tank no. 2 quantity probe short circuited at 

46:40:08. 

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. 

(2) 

22. Findings 

a, The crew acknowledged Mission Control's request to turn on 

the tank fans at 55:53:06. 

b. Spacecraft current increased by 1 ampere at 55:53:19. 

c. The oxygen tank no. 1 pressure decreased 8 psi at 55:53:19 

due to normal destratification. 

Determination 

Tne fans in oxygen tank no. 1 were turned on and began rotating 

at 55:53:19. 

23. Findings 

a, Spacecraft current increased by 1-1/2 amperes and ac bus 2 

voltage decreased 0.6 volt at 55:53:20. 

Stabilization and Control System (SCS) gimbal command telemetry channels, wnich are sensitive indicators of electrical 

transients associated with switching on or off of certain 

spazecraft electrical loads, showed a negative initial transient during oxygen tank no. 2 fan turnon cycles and a positilie initial transient during oxygen tank no. 2 fan turnoff 

b. 

5-20 

41-591 0 - 70 - 17 
cycles during the Apollo 13 mission. A 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) 

(2) 

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

It cannot 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. Finding 

An 11.1-amp spike in fuel cell 3 current and a momentary 

1.2-volt decrease were measured in ac bus 2 at 55:53:23. 

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. 

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. 

(2) 

25. Findings 

a. A momentary 11-volt decrease in ac bus 2 voltage was measured at 55:53:38. 

b. A 22.9-amp spike in fuel cell 3 current was measured at 

55: 53 :41. 

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. 

c. 

Determination 

Two short circuits occurred in the oxygen tank no, 2 fan circuits between 55:53:38 and 55:53:41 which resulted in either 

blown fuses or opened wiring, and the second fan ceased to 

function. 

5-21 
255 

26. Finding 

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

954 psia between 55:53:36 and 55:54:00. 

constant for about 15 seconds and then rose again from 954 to 

1008 psia, beginning at 55:54:15 and ending at 55:54:45. 

De terminations 

It then remained nearly 

(1) 

(2) 

An abnormal pressure rise occurred in oxygen tank no. 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. 

27. Findings 

a. The pressure relief valve was designed to be fully open at 

about 1000 psi. 

b. Oxygen tank no. 2 telemetry showed a pressure drop from 

1008 psia at 55:54:45 to 996 psia at 55:54:53, at which time 

telemetry data were lost. 

Determination 

This drop resulted from the normal operation of the pressure 

relief valve as 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 CM 

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 was alerted to the pressure rise. 

b. 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 

master alarm from being generated when any other parameter in 

the same system becomes out-of-tolerance, 

c. The low-pressure trip level of the master cadtion and warning 

system for the cryogenic storage system is only 1 psi below 

the specified lower limit of the pressure switch which controls the tank heaters. A small imbalance in hydrogen tank 

5-22 
25G 

pressures or a shift in transducer or switch calibration can 

cause the master 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 way 

to ascertain that the light did, in fact, come on. If ii 

did come on, Mission Control did not observe it. 

De terminations 

(1) If the pressure switch setting and master caution and warning 

trip levels were separated by a greater pressure differencial, 

there wo,Jld be less likelihood of unnecessary master alarms. 

With the present master caucion and warning system, a spacecraft problem can go unnoticed becaJse of the presence of a 

previous out-of-tolerance condition in the same subsystem. I 

Although a master alarm at 55:54:29 or observance of a limit 

serse light in Mission Control could have alerted the crew 

or Nissior Control in sufficient time to detect the pressure 

rise in oxygen tank no. 2. no action could have been taken 

at that time to prevent the tank failure. However, the information coEld have been helpful to Mission Control and the 

crew in diagnosis of spacecraft malfunctions. 

The limit sense system in Mission Control can be modified tc 

zonstitute a more positive backup warning syster.. 

(2) 

(3) 

(4) 

29. Fin3ing 

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

beginning at j5:54:3i sensed by a single sensor which measured 

local teaperature. This sensor indicated off-scale low at 

55: 54: 5:. 

Determinations 

(1) An abnormal and sudden temperature rise occurred in oxygen 

tank no. 2 at approximately 55:54:31. 

The temperature was a local value which rose when combustion 

had progressed to the vicinity of the sensor. (2) 

(3) The tercperature sensor failed at 5j:54:53. 

5-23 
257 

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 75.3 percent at 

55:54:32, and (3) quantity was off-scale high at 55:54:51 and 

later became erratic. 

Determinations 

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

55:54:50 may represent valid measurements. 

(2) Immediately preceding and following this time period, the 

indications were caused by electrical faults. 

31. Findings 

a. At about 55:54:53, or about half a second before telemetry 

loss, the body-mounted linear accelerometers in the command 

module, which are sampled at 100 times per second, began 

indicating spacecraft motions. I'hese disturbances were 

erratic, but reached peak values of 1.17g, 0.659, and 0.659 

in the X, Y, and 2 directions, respectively, about 13 milliseconds before data loss. 

b. The body-mounted roll, pitch, and yaw rate gyros showed lowlevel activity for 1/4 second beginning at 55:54:53.220. 

c. The integrating accelerometers indicated that a velocity 

increment of approximately 0.5 fps was imparted to the spacecraft between 55:54:53 and 55:54:55. 

Doppler tracking data measured an incremental velocity component of 0.26 fps along a line from the Earth to the spacecraft at approximately 55:54:55. 

d. 

e. The crew heard a loud "bang" at about this time. 

f. Telemetry data were lost between approximately 55:54:53 and 

55:54:55 and the spacecraft switched from the narrow-beam 

antenna to the wide-beam antenna. 

Crew observations and photographs showed the bay 4 panel to be missing and the high-gain antenna to be damaged. 

g. 

5-24 
258 

De t e rm inat'i ons 

(1) The spacecraft was subjected to abnormal forces at approximately 55:54:53. 

from failure and venting of the oxygen tank no. 2 system and 

subsequent separation and ejection of the bay 4 panel. 

The high-gain antenna was damaged either by the panel or a 

section thereof from bay 4 at'the time of panel separation. 

These disturbances were reactions resulting 

(2) 

32. Finding 

Temperature sensors in bay 3, bay 4, and the central column of 

the SM indicated abnormal increases following reacquisition of 

data at 55:54:55. 

Determination 

Heating took place in the SM at approximately the time of panel 

separation. 

33. Findings 

a. The telemetered nitrogen pressure in fuel cell 1 was offscale low at reacquisition of data at 55:54:55. 

b. Fuel cell 1 continued to operate for about 3 minutes past 

this time. 

c. The wiring to the nitrogen sensor passes along the top of 

the shelf which 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. 

The failure was probably caused by physical damage to the 

sensor wiring or shock. 

This is the only known instrumentation failure outside the 

oxygen system at that time. 

(2) 

(3) 

34. Finding 

Oxygen tank no. 1 pressure decreased rapidly from 879 psia to 

782 psia at approximately 55:54:54 and then began to decrease 

more slowly at 55:54:56. 

5-25 
259 

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 beginning 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 supply valves were 

cut off. 

The crew reported at 55:57:44 that five valves in the reaction 

control system (RCS) were 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. 

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

b. 

c. 

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. 

MISSION EVENTS AFTER ACCIDENT 

36. Findings 

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 was not directly noticed. 

However, the Guidance Officer did note and report a "hardware 

restart" of the spacecraft computer. 

followed by the crew's report of a problem. 

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: 

This was quickly 

b. 

5-26 
Oxygen tank no. 1: Pressure: Several hundred psi below 

normal 

&antity: Normal 

Temperature : Xormal 

Oxygen cank no. 2: Pressure: Off-scale low 

Quantity: Off-scale high 

Temperature: Off-scale high 

c. The nitrogen pressure in fuel Cell 1 indicated zero, which was 

incompatible with the nydrogen and oxygen pressures in this 

fuel cell, which were normal. 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. 

d. Neither the crew nor Mission Control was aware at the time 

that oxygen tank no. 2 pressure had risen abnormally just 

before the data loss. 

e. The flight controllers believed that a probable cause of 

these indicacions could have been a cryogenic storage system 

instrimentation failure, and began pursuing this line of in- 

.res tigation. 

Determination 

Under Lhese conditions it was reasonable to suspect a cryogenic 

storage system instrumentation problem, and to attempt to verify 

the readings before taking any action. The fact that the oxygen 

tank no. 2 quantity measurement was known to have 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 were less than 0.1 lb/hr. These were unusually 

low readings for the current being drawn. 

Fuel cells 1 and 3 failed at about 3 minutes after the data 

loss. 

b. 

5-27 
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 was 

initially believed to be a cryogenic system instrumentation 

problem to the electrical power system. 

Near-zero oxygen flow to fuel cells 1 and 3 was noted after 

the main bus B failure, but this was consistent with no power 

output from the fuel cells. 

d. 

e. The flight controllers believed that the fuel cells could 

have been disconnected from the busses and directed the crew 

to connect fuel cell 1 to dc main bus A 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. 

Determinations 

(1) Under these conditions it was logical for the flight controllers to attempt to regain power to the busses since the 

fuel cells might have 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 between a disconnected 

fuel cell and a failed one. 

If the crew had been aware of the reactant valve closure, 

they could have opened them before the fuel cells were starved 

of oxygen. This would have simplified subsequent actions. 

(2) 

38. Finding 

?he fuel cell reactant'valve talkback indicators in the spacecraft 40 not indicate closed unless both the hydrogen and oxygen 

valves are closed. 

Determinations 

(1) If these ialkbacks were designed so that either a hydrogen 

or oxygen valve closure would indicate "barberpole ," the 

Apollo 13 crew could possibly have acted in time to delay 

the failure of fie1 cells 1 and 3, although they would never 

theless have failed when oxygen tank no, 1 ceased to supply 

oxygen. 

5-28 
262 

(2) The ultimate outcome would not have been changed, but had the 

fuel cells not failed, Mission Control and the crew would not 

have had to contend with the failure of dc main bus B and ac 

bus 2 or attitude control problems while trying to evaluate 

the situation. 

Reaction Control System 

39. Findings 

a. 

b. 

C. 

d. 

e. 

The crew reported the talkback indicators for the helium 

isolation valves in the SM RCS quads B and D indicated closed 

shortly after the dc main bus B failure. 

pressurization valves for quads A and C also were reported 

closed. 

The SM RCS quad D propellant tank pressures decreased until 

shortly after the crew was requested to confirm that the 

helium isolation valves were opened by the crew. 

During the l-1/2-hour period following the accident, Mission 

Control noted that SM RCS quad C propellant was not being 

used, although numerous firing signals were being sent to it. 

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. 

During the l-1/2-hour period immediately following the 

accident, Mission Control advised the crew which SM RCS 

thrusters to power and which ones to unpower. 

The secondary fuel 

Determinations 

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

the accident: 

Helium isolation valves 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 same shock. 

Mission Control correctly determined the status of the RCS 

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

(3) 

5-29 
263 

Management of Electrical System 

40. Findings 

a. After fuel cell 1 failed, the total dc main bus A load was 

placed on fuel cell 2 and the voltage dropped to approximately 25 volts, causing a caution and warning indication 

and a master alarm. 

b. After determining the fuel cell 2 could not supply enough 

power to dc main bus A to maintain adequate voltage, 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 reduce the electrical 

load on dc main bus A by following the emergency powerdown 

checklist contained in the onboard Flight Data File. 

d. When the power requirements were sufficiently reduced so that 

the one remaining fuel cell could maintain adequate bus 

voltage, Mission Control directed the crew to take the entry 

battery off line. 

e. Mission Control then directed the crew to charge this battery 

in order to get as much energy back into it as possible, 

before the inevitable loss of the one functioning fuel cell. 

Determinations 

(1) Emergency use of the entry battery helped prevent potential 

loss of dc main bus A, which could have led to loss of communications between spacecraft and ground and other vital CM 

functions. 

Available emergency powerdown lists facilitated rapid reduction of loads on the fuel cell and batteries. 

(2) 

Attempts to Restore Oxygen Pressure 

41. Findings 

a. After determining that the CM problems were not due to instrumentation malfunctions, and after temporarily securing 

a stable electrical system configuration, Mission Control 

sought to improve oxygen pressures by energizing the fan 

and heater circuits in both oxygen tanks, 

5-30 
264 

b. '&en these procedures failed to arrest the oxygen loss, 

Mission Control directed the crew to shut down fuel cells 1 

and 3 by closing the hydrogen and oxygen flow valves. 

Determinations 

(1) L'nder more normal conditions oxygen pressure might have been 

increased by turning on heaters and fans in tne oxygen tanks; 

no oLher known actions had such a possibility. 

There was a possibility that oxygen was leaking downstream 

of the valves; had this been true, closing of the valves 

might nave preserved the remaining oxygen in oxygen tank 

no. 1. 

(2) 

Lunar Module Activation 

42. Findirgs 

a. With imminent loss of oxygen from oxygen tanks no. 1 and 

no. 2, and failing electrical power in the CM, it was 

necessary to use the lunar module (LM) as a "lifeboat" for 

the return to Earth. 

b. Plission Control and the crew delayed LM activation until 

about 15 minutes before Lhe SM oxygen supply was depleted. 

c, There were three different iM activation checklists contained 

in the Flight Data File for normal and contingency situations; 

noweflier, none of tnese was appropriate for the existing situation. It was necessary to activate the LM as rapidly as 

possible to conserve J3 consumables and CM reentry batteries 

to the maximum extent possible. 

d. Mission Control modified the normal LM activation checklist 

and referred tne crew to specific pages and instructions. 

This bypassed unnecessary steps and reduced the activation 

tine to less than an hour. 

The LM inertial platform was aligned during an onboard checklist procedure which manually transferred the CM alignment to 

the LM. 

e. 

5-31 
265 

Dete rminations 

(1) Initiation of LM activation was not undertaken sooner because 

the crew was properly more concerned with attempts to conserve 

remaining SM oxygen. 

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 LM. 

(2) 

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 

girbal lock condition. 

The Command Module Pilot (CMP) stated that he was not concerned because he could have quickly reestablished direct 

manual attitude control if it became necessary. 

c. 

Determination 

This situation was not hazardous to the crew because had gimbal 

locK actually occurred, sufficient time was available to reestablish an attitude reference. 

44. Findings 

a. L14 flight controllers were on duty in Mission Control at tne 

time of the accident in support of the scheduled crew entry 

inzo the LM. 

5. If the accident had occurred at some other time during the 

translunar coast phase, LM system specialists would not have 

Seen on duty, and it would have taken at least 30 minutes to 

get a fully manned team in Mission Control. 

Determination 

Although IN flight controllers were not recuired until more than 

an hour after the accident, it was beneficial for them to be 

present as the problem developed. 

5-32 
266 

LM Consumables Management 

45. Findings 

a. The LM was designed to support two men on a 2-day expedition 

to the lunar surface. Mission 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, 

b.. An emergency powerdown checklist was available in the Flight 

Data File on board the LM. Minor revisions were made to the 

list to reduce electrical energy requirements to about 

20 percent of normal operational values with a corresponding 

reduction in usage of coolant loop water, 

Mission Control determined that this maximum powerdown could 

be delayed until after 80 hours ground elapsed time, allowing 

the LM primary guidance and navigation system to be kept 

powered up for the second abort maneuver. 

c. 

d. Mission Control developed contingency plans for further reduction of LM power for use in case an LM battery problem 

developed, Procedures for use of CM water in the LM also 

were developed for use if needed. 

e. Toward the end of the mission, sufficient consumable margins 

existed to allow usage rates to be increased above earlier 

planned levels. This was done. 

hben the LM was jettisoned at 141:30 the approximate remaining 

margins were: 

Electrical power 4-1/2 hours 

Water 5-1/2 hours 

Oxygen 124 hours 

f. 

Determinations 

(1) Earlier contingency plans and available checklists were 

adequate to extend life support capability of the LM well 

beyond its normal intended capability. 

Mission Control maintained the flexibility of being able to 

further increase the LM consumables margins. 

(2) 

5-33 
Modification of LM Carbon Dioxide Removal System 

46. Findings 

a. The lithium hydroxide (LiOH) cartridges, which remove water 

and carbon dioxide from the LM cabin atmosphere, would have 

become ineffective due to saturation at about 100 hours. 

b. Mission rules set maximum allowable carbon dioxide partial 

pressure at 7.5mm Hg. LiOH cartridges are normally changed 

before cabin atmosphere carbon dioxide partial pressure 

reaches this value. 

c. Manned Spacecraft Center engineers devised and checked out a 

procedure for using the CM LiOH cannisters to achieve carbon 

dioxide removal. Instructions were given on how to build a 

modified cartridge container using materials in the spacecraft . 

d. The crew made the modification at 93 hours, and carbon 

dioxide partial pressure in the LM dropped rapidly from 

7.5m Hg to 0.lmm Hg. 

e. Mission 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 Earthreturn trip. 

Determination 

The Manned Spacecraft Center succeeded in improvising and checking 

out a modification to the filter system which maintained carbon 

dioxide concentration well within safe tolerances. 

LM Anomaly 

47. Findings 

a. During the time interval between 97:13:53 and 97:13:55. IM 

descent battery current measurements on telemetry showed a 

rapid increase from values of no more than 3 amperes per 

battery to values in excess of 30 amperes per battery. The 

exact value in one battery cannot be determined because the 

measurement for battery 2 was off-scale high at 60 amperes. 

5-34 
268 

b. At about that time the Lunar Module Pilot (LMP) heard a 

"thump" from the vicinity of the LM descent stage. 

men the LMP looked out the LM right-hand window, he observed 

a venting of small particles from the general area where the 

LM descent batteries 1 and 2 are located. This venting continued for a few minutes. 

c. 

d. Prior to 97:13 the battery load-sharing among the four 

batteries had been equal, hut immediately after the battery 

currents returned to nominal, batteries 1 and 2 supplied 9 

of the 11 amperes total. By 97:23 the load-sharing had returned to equal. 

e. There was no electrical interface between the LM and the CSM 

at this time. 

f. An MSC investigation of the anomaly is in progress, 

3eter3ina:iors 

(1) An anomalous incident occurred in the LJ4 electrical system 

at abou: 97:13:53 which appeared to be a short circLit. 

The th'mp and the venting were related to this anomaly. 

Tqe apparent short circdit cleared itself 

This anomaly was not directly related to the CSM or to the 

accident. 

?is anomaly represents a potentially serious electrical 

problem. 

(2) 

(3) 

(4) 

(5) 

CM aattery Recnarging 

48. Findings 

a. About one half of the electrical capacity of reentry 

battery A (20 of 40 amp-tours) was used during emergency 

conditions following the accident. A small part of the 

capacity of reentry battery B was used in checking out cc 

main bus E at 95 hours. 

catterles limited the amount of time the CM could operace 

after separation from the LM. 

Tile reduced charge remaining in the 

5-35 
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 CM batteries A and B. This procedure used 

the electrical umbilical between the LM and the CM which 

normally carried electrical energy from the CM to the LM. 

The procedure was nonstandard and was not included in checklists. 

d. The procedure was initiated at 112 hours and CM batteries A 

and B were fully recharged by 128 hours. 

Determination 

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 provide an extra 

margin of safety for the reentry operation. 

Trajectory Charges For Safe Return to Earth 

49, Findings 

a. After the accident, iz became apparent that the lunar landing 

could noL be accomplished and that the spacecraft trajectory 

must be altered for a retxm to Earth. 

b. A.t the time of the' accident, the spacecraft trajectory was 

one which would have returned it to the vicinity of the Earth, 

cut it would have been left in orbit about the Earth rather 

than reencering for a safe splashdown. 

c. To return the spacecraft to Earth, the following midcourse 

corrections were made: 

A 38-fps correction at 61:30, using the LM descent propulsion 

system (LIPS), required to return the spaascraft 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. 

A 7.8-fps burn at 105:18 using the DPS engine to lower Earth 

perigee fron 87 miles to 21 miles. 

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270 

A 3.2-fps correction at 137:40 using LM RCS thrusters, to 

assure that the CM would reenter the Earth's atmosphere at 

the center of its corridor. 

All course corrections were executed with expected accuracy 

and the CM reentered the Earth's atmosphere at 142:40 to 

return the crew safely at 142:54, near the prime recovery 

ship. 

Without the CM guidance and navigation system, the crew could 

not navigate or compute return-to-Earth maneuver target parameters, 

Determinations 

d. 

e. 

(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 SM; it put the spacecraft on a 

trajectory, within a few hours after the accident, which had 

the best chance for a safe return to Earth; it placed splashdown 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 

open an option to further reduce the flight time. 

fiission Control trajectory planning and maneuver targeting 

were essential for the safe return of the crew. (2) 

Entry Procedures and Checklists 

50. Findings 

a. Preparation for reentry required nonstandard procedures because of the lack of SM oxygen and electrical power supplies. 

b. The SM RCS engines normally provide separation between the 

SM and the CM by continuing to fire after separation. 

c. Apollo 13 SM RCS engines could not continue to fire after 

separation because of the earlier failure of the fuel cells. 

d. The CM guidance and navigation system was powered down due to 

the accident. The LM guidance and navigation system had also 

been powered down to conserve electrical energy and water. A 

spacecraft inertial attitude reference had to be established 

prior to reentry. 

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271 

e. 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 CM batteries, LM jettison was delayed as long as practical. The LM batteries were used to 

supply part of the power necessary for CM activation. 

The procedures for accomplishing the final course correction 

and the reentry preparation were developed by operations 

support personnel under the direction of Mission Control. 

g. 

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 simulators at MSC and 

KSC by members of the backup crew. 

The procedures were read to the crew about 24 hours prior to 

reentry, allowing the crew time to study and rehearse them. 

j. Trajectory evaluations of contingency conditions for LM and 

i. 

SM separation were conducted and documented prior to the 

mission by mission-planning personnel at MSC. 

k. Most of the steps taken were extracted from other procedures 

which had been 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 carefully developed, simulated, and 

practiced over a lonp training period. 

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

51. Findings 

a. During the reentry preparations, after SM jettison, there was 

a half-hour period of very poor communications with the CM 

due to the spacecraft being in a poor attitude with the LM 

present. 

b. This condition was not recognized by the crew or by Mission 

Control. 

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272 

Determination 

Some of the reentry preparations were unnecessarily prolonged by 

the poor communications, but since the reentry preparation timeline was not crowded, 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 LM roll 

attitude in preparation for LM jettison. This attitude put 

the C?.! very close to gimbal lock which, had it occurred, would 

have lost the inertial attitude reference essential for an 

automatic guidance system control of reentry. 

b. If gimbal lock had occurred, a less accurate but adequate 

attitude reference could have been reestablished prior to 

reentry. 

Determination 

The most significant consequence of losing the attitude reference 

in this situation would have been the subsequent impact on the 

remaining reentry preparation timeline. In taking the time to 

reestablish this reference, less time would have been available 

to accomplish the rest of the necessary procedures. The occurrence of gimbal lock in itself would not have significantly increased the crew hazard. 

5-39 
PART 4. RECOMMENDATIONS 

1. 

modified to: 

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 fire in 

the tank. 

b. Minimize the use of Teflon, aluminum, 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. ';re warning systems on board the Apollo spacecraft and in the Mission 

Control Cexter should be carefully reviewed and modified where appropriate, 

wi-h 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 caJtion and warning system logic to prevent an outof-limits alarm from blocking another alarm $when a second quantity in the 

same s,Jbsystem goes out of limits. 

c. Establishing a secon~ level of limlt sensing in Mission Control 

on critical quanti-ies with a visual or audible alarm which cannot be 

eas ilj. overlooked I 

a. Pro.iiding independent talkback indicators for each of the six 

fnel cell reactar.: valves plus a master alarm when any valve closes. 

4. Consumables and emergency equipment in the M and the 04 should be reviewed to determine whe:her steps s?.ould be taken to enhance their potential for use in a "lifeboat" mode. 

5. L!e Manned Spacecraft Center should complete the special tests and 

analyses now underway in order to understand more completely the details 

of the Apollo I3 accident. In addition, the lunar module power system 

anomalies should receive careful attention. Other NASA Centers should 

Zontinue their support to MSC in the areas of analysis and test. 

5-40 
274 

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 equipnent, including those which have previously been corrected or explained. Furthermore, critical decisions involving 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. 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 information developed in this investigation. 

8. 

ignition, and combustion in strong oxidizers at various g levels; and on 

the characteristics of supercritical fluids. Where appropriate, new NASA 

design standards should be developed. 

9. The Manned Spacecraft Center should reassess all Apollo spacecraft 

subsystems, and the engineering organizations responsible for them at 

MSC 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. here 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 tes'c, and operational experience. 

NASA should conduct additional research on materials compatibility, 

0 

5-41