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REPORT OF

APOLLO 13 REVIEW BOARD

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NATIONAL AERONAUTICS

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AND SPACE

REPRODUCEDBY

NATIONAL TECHNICAL

INFORMATION SERVICE

U.S. DEPARTMENTOFCOMMERCE

SPRINGFIELD,VA. 22161

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ADMINISTRATION
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

APOLLO 13 REVIEW BOARD

June 15, 1970

The Honorable Thomas O. Paine

Administrator

National Aeronautics and Space Administration

Washington, D.C. 20546

Dear Dr. 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 recommendations are necessary. In

addition, we will stand ready to reconvene at your request.

Sincerely yours,

Edgar M. Cortright

Chairman
°

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ii
PREFACE

The Apollo 13 accident, which aborted man's third mission to explore

the surface of the Moon, is a harsh reminder of the immensedifficulty

of this undertaking.

The total Apollo system of ground complexes, launch vehicle, and

spacecraft constitutes the most ambitious and demandingengineering

development ever undertaken by man. For these missions to succeed, both

menand equipment must perform to near perfection. That this system has

already resulted in two successful lunar surface explorations is a tribute

to those menand womenwho conceived, designed, built, and flew it.

Perfection is not only difficult to achieve, but difficult to maintain. The imperfection in Apollo 13 constituted a near disaster, averted

only by outstanding performance on the part of the crew and the ground

control team which supported them.

The Apollo 13 Review Board was charged with the responsibilities

of reviewing the circumstances surrounding the accident, of establishing

the probable causes of the accident, of assessing the effectiveness of

flight recovery actions, of reporting these findings, and of developing

recommendations for corrective or other actions. The Board has made

every effort to carry out its assignment in a thorough, objective, and

impartial manner. In doing so, the Board madeeffective use of the

failure analyses and corrective action studies carried out by the Manned

Spacecraft Center and was very impressed with the dedication and objectivity of this effort.

The Board feels that the nature of the Apollo 13 equipment failure

holds important lessons which, when applied to future missions, will

contribute to the safety and effectiveness of mannedspace flight.

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ladder

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LM descent

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Service module

Apollo 13 space vehicle configuration.

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control system

radiator

steerable

high gain antenna

extension

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CSM in ground test with bay 4 panel removed.

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Inflight photograph of service module showing damageto bay 4.

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viii
TABLEOFCONTENTS

LETTEROFTRANSMITTAL

PREFACE............................

TABLEOFCONTENTS.......................

Page

iii

ix

CHAPTERi - AUTHORITIES

Memorandum,April 17, 1970, from Administrator and

Deputy Administrator to Mr. Edgar M. Cortright ......

Memorandum,April 21, 1970, from Administrator and

Deputy Administrator to Mr. Edgar M. Cortright ......

Memorandum,April 20, 1970, from Administrator and

Deputy Administrator to Dr. Charles D. Harrington,

Chairman, Aerospace Safety Advisory Panel .........

Memorandum,April 20, 1970, from Administrator to

Mr. Dale D. Myers, Associate Administrator for

MannedSpaceFlight ....................

NASAManagementInstruction 8621.1, Subject:

Mission Failure Investigation Policy and Procedures,

April 14, 1966 ......................

NASAManagementInstruction 1156.14, Subject:

Aerospace Safety Advisory Panel, December7, 1967 .....

CHAPTER2 - BOARDHISTORYANDPROCEDURES

Part i. Summaryof Board History and Procedures • • • . . ,

Part 2. Biography of Board Members, Observers, and

Panel Chairmen

• . • . . • , • . , • • • , • . • , •

Part 3. Board Organization and General Assignments

for Board Panels

• • • • , . • • • . , • • • , • • •

Part 4. Summary of Board Activities

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1-6

1-7

i-8

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2-1

2-3
Page

CHAPTER 3 - DESCRIPTION OF APOLLO 13 SPACE VEHICLE AND MISSION

Part i. Apollo/Saturn V Space Vehicle .......... 3-2

Part 2. Apollo 13 Mission Description .......... 3-26

CHAPTER 4 - REVIEW AND ANALYSIS OF APOLLO 13 ACCIDENT

Part i. Introduction .................. 4-1

Part 2. Oxygen Tank No. History ............. 4-2

Part 3. Apollo 13 Flight ................ 4-25

Part 4. Surm_ary Analysis of the Accident ........ 4-36

Part 5. Apollo 13 Recovery ............... 4-44

CHAPTER 5 - FINDINGS, DETERMINATIONS, AND RECOMMENDATIONS

Part i. Introduction ................. 5-1

Part 2. Assessment of Accident ............. 5-5

Part 3. Supporting Considerations ............ 5-12

Part 4. Recommendations ................. 5-40

APPENDIX A - BASELINE DATA: APOLLO 13 FLIGHT SYSTEMS AND OPEBATIONS

APPENDIX B - REPORT OF MISSION EVENTS PANEL

APPENDIX C - REPORT OF MANUFAC_JRING AND TEST PANEL

APPENDIX D - REPORT OF DESIGN PANEL

APPENDIX E - REPORT OF PROJECT _G_NT PANEL

APPENDIX F - SPECIAL TESTS AND ANALYSES

APPENDIX G - BOARD ADMINISTRATIVE PRODECURES

APPENDIX H - BOkRD RELEASES AND PRESS STATEMEntS

II
CHAPTER i

AUTHORITIES

l-O

N N- N _ L E. E L l_ li
°
_iBI l_- _

OFFICE OF THE ADMINISTRATOR

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

WASHINGTON, D.C. 20546

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

i. 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 its space and aeronautical activities and to take appropriate

corrective actions as a result of the findings and recommendations."

2. Because of the serious nature of the accident of the Apollo 13 spacecraft 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:

(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.

i-i
(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.

(e) 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 recommendations 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.

i-2
7. All elements of NASA will cooperate with the Board and provide full

support within their areas of responsibility.

George M. Low

Deputy Administrator T.O. Paine

Administrator

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i:< ̧

OFFICE OF THE ADMINISTRATOR

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

WASHINGTON, D.C. 20546

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,

Hqs., USAF)

Mr. Vincent L. Johnson (Deputy Associate Administrator-Engineering,

Office of Space Science and Applications)

Mr. Milton Klein (Manager, AEC-NASA Space Nuclear Propulsion Office)

Dr. Hans M. Mark (Director, Ames Research Center)

Counsel:

Mr. George Malley (Chief Counsel, Langley Research Center)

0MSF 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)
Dr. CharlesD. Harrington (Chairman,NASAAerospaceSafety

AdvisoryPanel)

Mr. I. I. Pinkel (Director, AerospaceSafety Researchand

DataInstitute, LewisResearchCenter)

Congressional Liaison:

Mr. Gerald J. 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. O. Paine

Administrator

1-5

Ll 11 L: L_ L L
OFFIC r OF" THE ADMINISTRATOR

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

WASHINGTON, D.C. 20546

April 20, 1970

TO

SUBJECT :

Attachment:

References:

Dr. Charles D. Harrington

Chairman, Aerospace Safety Advisory Panel

Review of Procedures and Findings of Apollo 13 Review Board

(a) Memorandum dated April 17, 1970, to Mr. Edgar M.

Cortright, subject: Establishment of Apollo 13

Review Board

(a) Section 6, National Aeronautics and Space Administration

Authorization Act, 1968

(b) NMI 1156.14 - Aerospace Safety Advisory Panel

i. 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 the 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. The final report and any interim reports of the Board will be made

available promptly to the Panel for its review.

5. The Panel is requested to report to us on the procedures and findings

of the Board at such times and in such form as you consider appropriate,

but no later than i0 days after the submission to us of the final report

of the Board.

George M. Low

Deputy Administrator

Enclosure

T. O. Paine

Administrator

cc: Mr. Edgar M. Cortright, Chairman, Apollo 13 Review Board

M/Mr. Dale Myers

i-6
OFFICE OF THE ADMINISTRATOR

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

WASHINGTON,D.C. 20546

April 20, 1970

TO

SUBJECT :

References:

Mr. Dale D. Myers

Associate Administrator for Manned Space Flight

Apollo 13 Review

(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

i. As indicated in paragraph 6 of Reference (a), you are directed to:

(a) 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.

(b) 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.

Deputy Administrator Administrator

cc: Mr. Edgar M. Cortright, Chairman, Apollo 13 Review Board

Mr. Charles D. Harrington, Chairman, Aerospace Safety Advisory Panel

I-7

L L E 12 L I: L L,: L.; n n Li n L ,
NMI 862 i. i

April 14, 1966

_ec_h,e dc#e

Management Instruction

SUBJECT: MISSION FAILURE INVESTIGATION POLICY AND PROCEDURES

Io P_P_E

This Instruction establishes the policy and procedures for investigating

and documenting the causes of all major mission failures which occur in the

conduct of NASA space and aeronautical activities.

2. APPLICABILITY

3.

5.

This Instruction is applicable to NASA Headquarters and field installations.

DEFINITION

For the purpose of this 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 Officlals-ln-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:

(l) 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 April 14, 1966

(2) Determining the causes or probable causes of all failures,

taking corrective or other actions, and submitting written

reports of such determinations and actions to the Deputy

Administrator.

Do When the Deputy Administrator decides to conduct an independent

investigation, he will:

(i) Establish a (name of project) Review Board, comprised of appropriate NASA officials;

(2) Define the specific responsibilities of each Board, encompassing

such tasks as:

(a) 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.

(b) Reviewing the findings during the course of investigations

with cognizant field installation and Headquarters officials.

(c) Recommending 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 nontechnical contributing factors.

(d) Developing recommendations for corrective and other actions,

based on all information available to the Board.

(e) Doc_nenting findings, determinations and recommendations

for corrective or other actions and submitting such documentation to the Deputy Administrator.

c. Procedures for implementing the Board's reco_mnendations shall be

determined by the Deputy Administrator.

6. CANCELLATION

NASA Management Manual Instruction 4-1-7 (T.S. 760), March 24, 1964.

DISTRIBUTION:

SDL I

Deputy Administrator

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N_ n>6._4

December 7, 1967

C'ffectiv¢dole

Management Instruction

SUBJECT: AEROSPACE SAFETY ADVISORY PANEL

i. PURPOSE

This Instruction sets forth the authority for, and the

duties, procedures, organization, and support of the

Aerospace Safety Advisory Panel.

2. AUTHORITY

The Aerospace Safety Advisory Panel (hereafter called the

"Panel") was established under Section 6 of the National

Aeronautics and Space Administration Authorization Act,

1968 (PL 90-67, 90th Congress, 81 Stat. 168, 170). Since

the Panel was established by statute, its formation and

use are not subject to the provisions of Executive Order

ll007 or of NMI 1150.2, except to the extent that such

provisions are made applicable to the Panel under tflis

Instruction•

3. DUTIES

aB The duties of the Panel are set forth in Section 6

of the National Aeronautics and Space Administration

Authorization Act, 1968, as follows:

"The Panel shall review safety studies and

operations plans referred to it and s_all

make reports 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."

b • Pursuant to carrying out its statutory duties, the

Panel will review, evaluate, and advise on all

elements of NASA's safety system, including

especially the industrial safety, systems safety,

1-10
NMI1156.14 December7, 1967

and public safety activities, and the managementof

these activities. These key elements of NASA's

safety system are identified and delineated as follows:

(1) 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 NASA'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, development, manufacture, test, operation, and maintenance. 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

design of new facilities.

(2)

(d) Establishment and implementation of safety

standards and procedures for operation of

program support and administrative aircraft.

Systems Safety. This element includes those

activities specifically organized to deal with the

potential hazards of complex R&D systems that

involve many highly specialized areas of technology. It places 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) Performance of systems safety analyses.

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ILl II ]J E L E L L l_ I_ L: A: _ n n m n L ' ,__
_cember 7, 1967 NMI i156.14

(3)

(4)

(d) Establishment and implementation of systems

safety plans.

Public Safety. This element includes those

activities which, on a continuing basis, provide

protection for the well being of people and

prevention of damage to property not involved in

_ASA'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 implementation, as required,

of emergency or catastrophe control plans.

Safety Management. This element includes both the

program and functional organizations of NASA 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 management

systems for planning, implementing, coordinating,

and controlling these activities. These management

systems include, but are not limited 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.

(c) The plans and procedures for accident/Incident

investigations, including those for the followup on corrective actions and the feedback of

accident/Incident information to other

involved or interested organizations.

(d) The analysis and dissemination of safety data.

1-12
NMI 1156.14 December 7, 1967

•

•

PROCEDURES

a• The Panel will function in an advisory capacity to the

Administrator, and, through him, to those organizational

elements responsible for management of the NASA safety

activities.

b, The Panel will be provided with all information required

to Uischarge its advisory responsibilities as they

pertain to both NASA and its contractors' safety

activities. This information will be made available

through the mechanism of appropriate reports, and by

means of in situ reviews of safety activities at the

various NASA and contractor sites, as deemed necessary

by the Panel and arranged through the Administrator.

The Panel will thus be enabled to examine and evaluate

not only the general status of the NASA safety system,

but also the key elements of the _lanned and on-going

activities in this system.

ORGANIZATION

a. [!embership

(l) The Panel will consist of a maximum of nine members,

who will be appointed by the Administrator.

Appointments will be for a term of six years,

except that, in order to provide continuity of

membership, one-third of the members appointed

originally to the Panel will be appointed for a

term of two years, one-third for a term of four

years, and one-third for a term of six years•

(2) Not more than four members of the Panel shall be

employees of NASA, nor shall such NASA members

constitute a majority of the composition of the

Panel at any given time.

(3) Compensation and travel allowances for Panel

members shall be as specified in Section 6 of the

NASA Authorizatfon Act, 1968.

b. Officers

(I) The Officers of the Panel shall be a Chairman and

a Vice Chairman, who shall be selected by the Panel

from their membership to serve for one-year terms.

(2) The Chairman, or Vice Chairman in his absence,

shall preside at all meetings of the Panel and shall

have the usual powers of a presiding officer.

1-13
December 7, 1967 NMI I156.14

Co

dQ

Committees

(z)

(2)

The Panel is authorized to establish special

committees, as necessary and as approved by the

Administrator, to carry out specified tasks within

the scope of duties cf the Panel.

All such cc_:_mittee activities will be considered

an inseparable extension of Panel activities, and

will be in accordance with all applicable procedures and regulations set forth in this

Instruction.

(3)

(4)

(5)

The Chairman of each special committee shall be a

member of the Aerospace Safety Advisory Panel. The

other committee members may or may not be members

of the Panel, as recommended by the Panel and

approved by the Administrator.

Appointment of Panel members to committees as

officers or members will be either for one year,

for the duration of their term as Panel members, or

for the lifetime of the committee, whichever is the

shortest. Appointments of non-Panel members to

committees will be for a period of one year or for

the lifetime of the committee, whichever is shorter.

Compensation and travel allowances for committee

members who are not members of the Panel'shall be

the same as for members of the Panel itself, except

that compensation for such committee members

appointed from outside the Federal Government shall

be at the rate prescribed by the Administrator for

comparable services.

Meetings

(l)

(2)

Regular meetings of the Panel will be held as often

as necessary and at least twice a_ar. One meeting

each year shall be an Annual Meeting. Business

conducted at this meeting will include selecting

the Chairman and the Fice Chairman of the Panel,

recommending new committees and committee 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

Chairman, by notice served personally upon or by

mall or telegraph to the usual address of each

member at least five days prior to the meeting.

1-14
-" NMI 1156.14 December 7, 1967

e.

f.

(3) Special meetings shall be called in the same

manner by the Chairman, upon the written request

of three members of the Panel.

(4)

(5)

(6)

If practicable, the object of a special meeting

should be sent in writing to all members, and if

possible a special meeting should be avoided by

obtaining the views of members by mail or otherwise,

both on the question requiring the meeting and on

the question of calling a special meeting.

All meetings of special committees will be called

by their respective qhairmen pursuant to and in

accordance with performing their specified tasks.

Minutes of all meetings of the Panel, and of special

committees established by the Panel, will be kept.

Such minutes shall, at a minimum, contain a record

of persons present, a description of matters 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

committee.

Reports and Records

(1) The Panel shall submit an annual report to the

Administrator.

(2) The Panel will submit to the Administrator reports

on all safety reviews and evaluations with comments

and recommendations as deemed appropriate by the

Panel.

(3) All records and files of the Panel, including

agendas, minutes of Panel and committee meetings,

studies, analyses, reports, or other data compilations or work papers, made available to or

prepared by or for the Panel, will be retained by

the Panel.

Avoidance of Conflicts of Interest

(l) Nongovernmental members of the Panel, and of

special committees establishea by the Panel, are

"Special Government Employees" within the meaning

of NHB 1900.2A, which sets forth guidance to NASA

Special Government Employees resarding the

avoidance of conflicts of interest and the

observance of ethical standards of conduct. A

1"19
_ceaber 7, i_? _ 1156.14

. SUPPORT

a.

be

(2)

copy of NHB 1900.2A and related NASA instructions

on conflicts of interest will be furnished to each

Panel or committee member at the time of his

appointment as a NASA consultant or expert.

Nongovernmental members of the Panel or a soecial

committee will submit a "NASA Special Government

Employees Confidential Statement of Employment

and Financial Interests" (NASA Form 1271) prior to

participating in the activities of the Panel or a

special committee.

A staff, to be comprised of full-time NASA employees,

shall be established to support the Panel. The members

of this staff will be fully responsive to direction from

the Chairman or t_e Fanel.

The director of this staff will serve as Executive

Secretary to the Panel. The Executive Secretary of the

Panel, in accordance with the specific instructions from

the Chairman of the Panel, shall:

(1) Administer the affairs of the Panel and have general

supervision of all arrangements for safety reviews

and evaluations, and other matters undertaken by

the Panel.

(2)

(3)

Insure that a written record is kept of all

transactions, and submit the same to the Panel for

approval at each subsequent meeting.

Insure that the same service is provided for all

special committees of the Panel.

trator

CFR Title i_, Chapter 5, Subpart 1209.5.

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CHAPTER2

BOARDHISTORYANDPROCEDURES

Z-O

PART i. SUMMARY OF BOARD HISTORY AND PROCEDURES

The Apollo 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 manned by a number

of specialists, thereby providing a nucleus of expertise for the review

activity. During the period of the Board's review activities, the

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, USAF, was also conducting its own analysis of the

accident on Apollo 13. Coordination between the Investigation Team

work and the Apollo 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
The Board Chairman established a series of administrative procedures

to guide the Board's activities. In addition, specific assignments of

responsibility were madeto 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 MannedSpacecraft Center

Investigation Team.

In general, the Board relied on MannedSpacecraft 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 MSCbut also included tests run at other

NASACenters. Membersof 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 NASADeputy Administrator

with oral progress reports. These reports summarizedthe 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 as 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. In addition, the MSCInvestigation Teamestablished a file of

data at MSC.

2-2
PART2. BIOGRAPNT_S OF BOARDMEMBERS_ OBSERVERS_ AND PANEL CHAIRMEN

CHAIRMAN OF THE APOLLO 13 REVIEW BOARD

EDGAR 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 i0 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. In 1960, he

became Assistant Director for Lunar and Planetary Programs and directed

the planning and implementation of such projects as Mariner, Ranger, and

Surveyor.

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 of 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."

He is a native of Hastings, Pennsylvania, and served as a U.S. Navy

officer in World War II. 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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MEMBERSOFTHEAPOLLO13 REVIEWBOARD

ROBERTF. ALLNUTT

NASAHeadquarters

Robert F. Allnutt_ 34, Assistant to the NASAAdministrator,

Washington, D. C., is a memberof the Apollo 15 Review Board.

Mr. Allnutt was namedto his present position this year. Prior to

that, he had been Assistant Administrator for Legislative Affairs since

1967.

He joined NASAin 1960 as a patent attorney at the Langley Research

Center, Hampton,Virginia. In 1961, he was transferred to NASAHeadquarters, Washington, D. C.

Mr. Allnutt served as Patent Counsel for CommunicationsSatellite

Corporation from January to September1965, whenhe returned to NASA

Headquarters as Assistant General Counsel for Patent Matters.

He is admitted to the practice of law in the District of Columbia

and the state of Virginia and is a memberof 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

Nell A. Armstrong, 39, NASA astronaut, is a member of the Apollo 13

Review Board.

Commander of the Apollo ll 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.

\
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. He performed the first successful docking of

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.

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

AIAAAstronautics 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
Dr. Clark began his career in 19h2 as an electronics engineer at

the Naval Research Laboratory, Washington, D.C. From 19h7 to 19h8 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 II. After the War, he was assigned to the 19th Air Force, the

lhth 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 Energy.

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 h951st Support Squadron, Eniwetok;

and the following year, he was reassigned to Kirtland AFB as Assistant

to the Group Commander and later as Air Commander of the h925th Test Group.

General Hedrick Joined the Special Systems Office, Air Force

Ballistics Division, Los Angeles, in 1960. He was named Commander of

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-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. He received

B.S. and M.S. degrees in physics from the University of Maryland.

General and Mrs. Hedrick are the parents of two sons.

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 13 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 II, 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.

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L E L L I: L L "
MILTON KLEIN

NASA Headquarters

Milton Klein, 46, Manager, Space Nuclear Propulsion Office, NASA

Headquarters, is a member of the Apollo 13 Review Board.

Mr. Klein has been in his present position since 1967. Prior to

that he had been Deputy Manager since 1960. The Space Nuclear Propulsion

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.

Mr. Klein became associated with atomic energy work in 19467 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.

Mr. Klein was born in St. Louis, Missouri. He served in the U.S.

Navy during World War II.

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 Department 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 _ research physicist at

the University's Lawrence Radiation Laboratory, Livermore, California,

2-8
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 ll years old. He became a naturalized U.S. citizen

in 1945.

Dr. and Mrs. Mark are the parents of two children.

COUNSEL TO THE APOLLO 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 Langley 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. His last assignment was commanding

officer of the U.S.S. Fentress.

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.

2-9
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. In his present assignment, he serves as general

manager of manned space flight.

Prior to his 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. He served

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.

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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iu I/ Ju E L L L h ' L E E ;
APOLLO 13 REVIEW BOARD OBSERVERS

WILLIAM A. ANDERS

National Aeronautics and Space Council

William A. Amders, 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 ll and backup command module pilot for

Apollo ll, 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. Anders 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

managementpositions, was appointed Vice President, Mallinckrodt Nuclear

Corporation and Vice President, Mallinckrodt Chemical Works.

In 1961, whenthe fuel material processing plant of Mallinckrodt

becamethe Chemicals Division of United Nuclear Corporation, Dr. Harrington

was namedVice President of that division.

He becameSenior Vice President, United Nuclear Corporation,

Centreville, Maryland, in 1963.

In 1965, Dr. Harrington was appointed President and General Manager,

Douglas United Nuclear, Inc. The companymanagesproduction reactors

and fuels fabrication facilities at Hanford, Washington, for the Atomic

Energy Commission.

He is the co-author of a book, "Uranium Production 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. IRVINGPINKEL

NASALewis Research Center

I. Irving Pinkel, 57, Director, Aerospace Safety Research and Data

Institute at the NASALewis 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 Governmentscientific 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. Whenthe Lewis Research Center wasbuilt in

1942, he transferred there.

2-12
He has been elected to Phi Beta Kappa, SigmaXi, honorary scientific

society, and Pi MuEpsilon, honorary mathematics fraternity. He is an

Ohio Professional Engineer, served on the former NACAsubcommittees on

Meteorological Problems, Icing Problems, Aircraft Fire Prevention and

Flight Safety, and is a memberof the NASAResearch and Technology Advisory Subcommitteeon 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 NACADistinguished Service Medal, and the NASASustained Superior

Performance Award.

He wasborn in Gloversville, NewYork, and was graduated from the

University of Pennsylvania.

Mr. and Mrs. Pinkel live in Fairview Park, Ohio. They are the

parents of two sons.

JAMESE. WILSON,JR.

Committeeon Science and Astronautics

United States Houseof Representatives

JamesE. 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. From 1961 to 1963, he was Director of Research and Development,

U.S. Naval Propellant Plant, Indian Head, Maryland. Mr. Wilson managed

the Polaris Program at Indian Headfrom 1956 to 1961.

From 1954 to 1956, Mr. Wilson served as an officer in the U.S. Army

Signal Corps. He wasa development engineer with E. I. DuPont, Wilmington,

Delaware, from 1953 to 1954.

Mr. Wilson is a memberof Phi SigmaAlpha, a National Honor Society;

American Institute of Chemical Engineers; American Chemical Society; and

American OrdnanceAssociation.

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.

APOLLO13 REVIEWBOARDPANELCHAIRMEN

SEYMOURC. HIMMEL

NASALewis Research Center

Dr. SeymourC. Himmel, Assistant Director for Rockets and Vehicles,

Lewis Research Center, Cleveland, Ohio, heads the Design Panel of the

Apollo 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 NASAExceptional Service Medal and the NASA

Group Achievement Award as managerof the AgenaProject Group. Dr. Himmel

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. Himmelhas a Bachelor of Mechanical Engineering degree from the

College of the City of NewYork and M.S. and Ph.D. degrees from Case

Institute of Technology.

Dr. and Mrs. Himmel live in Lakewood, Ohio.

EDWINC. KILGORE

NASALangley Research Center

Edwin C. Kilgore, 47, Deputy Chief, Engineering and Technical Services, Langley Research Center, Hampton,Virginia, heads the Project

ManagementPanel of the Apollo 15 Review Board.

Mr. Kilgore joined the Langley science staff in 1944 and served in

a variety of technical and managementpositions until promotion to his

present position in 1968.

He has received the Honorary GroupAchievement Award for his role

in achieving a record of 97 consecutive successes for solid propellant

rocket motors and the NASA-LunarOrbiter Project GroupAchievement Award

for outstanding performance. He is a memberof Pi Tau Sigma, honorary

mechanical engineering society.

2-1h
Mr. Kilgore was born in Coeburn, Virginia. He was graduated from

Virginia Polytechnic Institute with a B.S. degree in mechanical engineering.

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, Minnesota. 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 II. 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

L L: E L L ' U. L: ig If 1_: L: L L_ L_
Mr. Smith was born in Piedmont, South Carolina, and received a B.S.

degree in electrical engineering 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

i,i I.l 12 L L L U L L_ L '_ "" _'
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
r

ro

i

(Do

OBSERVERS

W. A. ANDERS

C. D. HARRINGTON

I. I. PINKEL

J. E. WILSON, JR.

t .................

APOLLO 13

REVIEW BOARD

E. M. CORTRIGHTo CHAIRMAN

R. F. ALLNUTT V.L. JOHNSON

N. A. ARMSTRONG M. KLEIN

J. F. CLARK H.M. MARK

W. R. HEDRICK

CONSULTANTS

W. D. ERICKSON

R. VAN OOLAH

I

I-in TECHNICAL SUPPORT

C, W. MATHEWS

I

COUNSEL

G. T. MALLEY

I

MISSION EVENTS I

F. B. SMITH

PRE-INCIDENT EVENTS

J. J. WILLIAMS

INCIDENT EVENTS

T. B. BALLARD

POST INCIDENT

EVENTS

M. P. FRANK

I

SECRETARIAT

E. P. SWIEDA

I

EDITORIAL I

GROUP

R. G. ROMATOWSKI

I

PUBLIC

AFFAIRS

B. M. DUFF

I

I LEG,SLAT,VE I

AFFAIRS

G.

J. MOSS NGHOFF

I

MANUFACTURING

& TEST

H. M. SCHURMEIER

FABRICATION & ]

ACCEPTANCE TESTING

E. F. BAEHR

SUBSYSTEM & I

SYSTEM TESTING

K. L. HEIMBURG

t RELIABILITY &

QUALITY ASSURANCE

B. T. MORRIS

I

DESIGN

S. C. HIMMEL

DESIGN I

EVALUATION

R. N. LINDLEY

W. F. BROWN, JR.

FAILURE MODES I

& MECHANISMS

W. R. LUCAS

ELECTRICAL

R. C. WELLS

RELATED SYSTEMS

J. F. SAUNDERS, JR.

I

PROJECT

MANAGEMENT

E. C. KILGORE

I

R. D. GINTER I

]

M. H. MEAD

J. B. WHITTEN

APPROVED

E. M. CORTIIGHT

APOLLO 15 REVIEW BOARD ORGANIZATION
GENERAL ASSIGNMENTS FOR BOARD PANELS

(AS DOCUMENTED IN THE BOARD'S ADMINISTRATIVE PROCEDURES)

Panel i - 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 Apollo 13 Review Board an 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 strategy.

Panel i Membership

Mr. F. B. Smith, Panel Chairman

Assistant Administrator for University Affairs

NASA Headquarters

Washington, D. C.

2-19

N N-IJ I: E L L '"
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. Nell Armstrong, Board Member and Panel Monitor

As tron aut

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 as 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 Testin6.- 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 numberunits 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 design 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 Panel 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 activity

shall also review and evaluate proposed design modifications, including

changes in operating procedures required by such modifications.

Failure Modes and Mechanisms.- This activity shall review the design

of the systems to ascertain the possible sources of failure and the m_mer

in which failures may occur. In this process, they shall attempt to

correlate 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 that 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 proposed design

modifications, including changes in operating procedures required by such

modi fi cati ons.

Related Systems .- This activity shall review the design of all

systems similar to that involved in the Apollo 13 incident with the view

to 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, Panel Chairman

Assistant Director for Rockets and Vehicles

Lewis Research Center

Cleveland, Ohio

Mr. William F. Brown, Jr.

Chief, Strength of Materials Branch

Materials and Structures Division

Administration Directorate

Lewis Research Center

Cleveland, Ohio

2-22

iu la lJ /_- /d L If L '
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

Huntsville, Alabama

Mr. J. F. Saunders, Jr.

Project Officer for Command and Service Module

Office of Manned Space Flight

NASA He adquarters

Washington, D. C.

Mr. Robert C. Wells

Head, Electric Flight Systems Section

Vehicles Branch

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 Management Panel

The Project Management Panel will undertake the following tasks:

I. 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 Apollo 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 managementlessons learned from the Apollo

13 mission from the standpoint of their applicability to subsequent

Apollo missions.

Tasks i and 2, above, should encompassboth 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 AdvancedResearch and Technology

NASAHeadquarters

Washington, D.C.

Merrill H. Mead

Chief of Programs and Resources Office

AmesResearch Center

Moffett Field, California

JamesB. Whitten

Assistant Chief, Aeronautical and Space Mechanics Division

Langley Research Center

Hampton,Virginia

Milton Klein, Board Memberand Panel Monitor

Manager, AEC-NASASpace 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

NASAAerospace Safety Advisory Panel

Washington, D.C.

2-24
I. Irving Pinkel

Director

Aerospace Safety Research and Data Institute

Lewis Research Center

Cleveland, Ohio

Mr. JamesE. Wilson

Technical Consultant to the Committeeon Science and Astronautics

United States Houseof Representatives

Washington, D.C.

Apollo 13 Review Board Support Staff

Brian M. Duff

Public Affairs Officer

MannedSpacecraft Center

Houston, Texas

Gerald J. Mossinghoff

Director of Congressional Liaison

NASAHeadquarters

Washington, D.C.

EdwardF. Parry

Counsel to Office of MannedSpaceFlight

NASAHeadquarters

Washington, D.C.

RaymondG. Romatowski

Deputy Assistant Director for Administration

Langley Research Center

Hampton,Virginia

Ernest P. Swieda

Deputy Chief, Skylab Program Control Office

KennedySpace Center, Florida

Consultants to the Board

Dr. WayneD. Erickson, Head

Aerothermochemistry Branch

Langley Research Center

Hampton,Virginia

Dr. Robert Van Dolah

Acting ResearchDirector

Safety Research Center

Bureau of Mines

Pittsburgh, Pennsylvania

2-25
MSCSupport to the Board

These persons were detailed by MSCto support the Apollo 13 Review

Board during its review activity at MSC. They are identified by MSC

position title.

Roy C. Aldridge

Assistant to the Director of Administration

Mary Chandler

Secretary

Jamie Moon

Technical Editor

Rex Cline

Technical Writer/Editor

Dorothy Newberry

Administrative Assistant

Evon Collins

Program Analyst

Lettie Reed

Editorial Assistant

Leroy Cotton

Equipment Specialist

Charlene Rogozinski

Secretary

MaureenCruz

Travel Clerk

Joanne Sanchez

Secretary

Janet Harris

Clerk Stenographer

Billie Schmidt

EmployeeDevelopmentSpecialist

Marjorie Harrison

Secretary

Frances Smith

Secretary

Phyllis Hayes

Secretary

George Sowers

ManagementPresentations Officer

William N. Henderson

ManagementAnalyst

Elaine Stemerick

Secretary

Sharon Laws

Secretary

Mary Thompson

Administrative Assistant

Carolyn Lisenbee

Secretary

Alvin C. Zuehlke

Electrical Engineer

Judy Miller

Secretary

2-26
PART 4. SUMMARY OF BOARD ACTIVITIES

APRIL 19, 1970

Chairman E. M. Cortright met with Langley officials to begin planning

the Apollo 13 Review Board approach. Tentative list of Panel Members and

other specialists were developed for consideration.

APRIL 20, 1970

Chairman Cortright met with the NASA Administrator, Deputy Administrator, and key NASA officials in Washington, D.C., to discuss Board

membership.

The Chairman met with NASA Office of Manned Space Flight top officials 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 support of

the Apollo 13 Board.

Chairman Cortright met with members of the Press to report 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
APRIL 22, 1970

The Board met with Colonel McDivitt's MSCInvestigation Teamto review the progress madeby MSCin identifying causes of the accident and

in developing an understanding of sequences and relationships between

known inflight events. In addition, MSCofficials briefed the Board on

MSCInvestigation Teamstructure and assignments.

The Board met with Panel i of the MSCInvestigation Teamfor detailed discussion of inflight events and consideration of early conclusions on implications of preliminary data analysis.

The Board held its second meeting to discuss MSCinvestigative

efforts and additional appointments of Panel specialists.

Board membersattended Panel i evening roundup of day's evaluation

activities, _nich included detailed discussions of specific studies,

data reductions, and support test activities already underway.

APRIL 23, 1970

The Apollo 13 Review Board established itself in proximity to the

MSCInvestigation Teamin Building 45, and arranged for all administrative 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 information.

Initial task assignments and responsibilities were madeto Board

Panels as guidance for detailed review work. Individual Board members

were assigned Panel overview responsibilities or other special tasks.

Administrative procedures were developed for Board activity, particularly to provide efficient interface with MSCpersonnel.

Board and Panel Membersagain met with MSCofficials to further review the sequence of events in the Apollo 13 mission and to examine early

hypotheses concerning causes 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 NASAspecialists in order to

bring Panels up to strength.

2-28
APRIL 24, 1970

Board Members, Panel Chairmen, and MSC officials reviewed additional

data analysis made by MSC and contractor personnel with particular emphasis upon the service module (SM) cryogenic system.

The 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 Ryker of NR on

NR's activities involving design, qualification, and tests of SM cryogenic oxygen tanks.

APRIL 25, 1970

The 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

L
progress in an intensive analysis of the Apollo 13 malfunction, including

a review of approved special tests. Oxygen tank, fuel cell components,

assemblies, and other 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 module EPS failure.

Additional Board specialists arrived at MSC and received detailed

briefings by MSC and Board personnel on selected aspects of the Apollo 13

data.

Panel Members received 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, !970

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 i. 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. yon 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 KSCduring the Apollo 13 countdowndemonstration test (CDDT).

Board and Panel personnel reviewed progress to date at a general

Board meeting involving all ReviewBoard 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 recommendations involving oxygen tank redesign.

The Board continued to review and examine oxygen tank ignition

sources and combustion propagation processes with specialists from MSC,

other NASACenters, and contractor personnel.

The Mission Events Panel continued to examine and record details of

all significant mission events as a basis for other Panel evaluations and

study.

Chairman Cortright convened two Board meetings to review Panel progress to date and to discuss work plans for the next several days.

The Project ManagementPanel visited North American Rockwell at

Downeyto review detailed procedures for acceptance tests, subcontractor

inspections, project documentation, and other managementinterface areas.

APRIL30, 1970

The Safety Advisory Panel continued discussions with Board Chairman

and MSCofficials on progress of total Apollo 13 review efforts.

Panel Membersreviewed 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 Managementpersonnel visited BeechAircraft Corporation to

review procedures used for assembly of cryogenic oxygen tanks and to discuss communication and information systems within the Apollo Program.
Panels continued to review detailed data in their respective areas.

MAYi_ 1970

Board and Panel personnel participated in a joint MSC/Apo!!o 13

Review Board status presentation to the NASADeputy Administrator. The

meeting covered all significant Apollo 13 findings and early conclusions

on the cause of the accident and appropriate remedial actions.

The MSCstaff briefed Board Memberson initial evaluations of proposed design changes in oxygen tank system.

Panel Memberscontinued to assess data accumulated from the Apollo 13

mission with particular emphasisupon the design and performance of electric power systems used in the service module.

Board Membersand Panel Chairmenreviewed specific test matrix being

proposed by Apo!lo 13 Review Board specialists covering most significant

unknownsinvolved in understanding failure mechanisms.

MAY2, 1970

Board Membersmet 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 ManagementPanel 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 into logical failure mode

patterns established by MSCand Board personnel.

MAY3, 1970

Chairman Cortright and Board Membersconducted 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 significant mission

events and explanatory data, including the support tests required to

clarify questions raised by events.

Panel Membersworked on individual analyses with particular attention

to developing requirements for additional test activity in support of tentative conclusions.

The Board agreed to strengthen its technical reviews of combustion

propagation and electrical design by adding specialists in these areas.

MAY4, 1970

The Design Panel continued its intensive review of the "shelf drop"

incident at NRinvolving the cryogenic oxygen flight tank used in

Apollo 13 in order to understand possible results of this event.

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.

MAY5, 1970

The Board met in General Session to discuss the scope and conduct of

support test activity, including careful documentation of test methods and

application of test results.

MSCpersonnel briefed Panel Memberson availability of additional

telemetry data in the MSCdata bank in order to insure Board consideration of all possible useful data.

Panels commencedinitial drafting of preliminary findings in specific

areas, including summarydescriptions of system performance during the

Apollo 13 flight.

The Board met with the MSCInvestigation Teamfor complete review of

the proposed test program.

2-33
. MAY6, 1970

Board Members,MSCpersonnel_ and Membersof NASA'sAerospace 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 MSCastronauts.

Panel Memberscontinued 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.

MAY7, 1970

The General Board Session reviewed complete analysis and test support

activities being conducted for the Board and MSCat various governmental

and contractor installations.

Board and Panel Membersmet to discuss Ameslaboratory tests concerning liquid oxygen combustion initiation energies required in the

cryogenic oxygen tank used in the Apollo 13 SM.

Panel i Membersreviewed 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.

MAY8, 1970

Dr. Robert Van Dolah, Bureau of Mines, joined the Board as a consultant on combustion propagation and reviewed Apollo 13 ReviewBoard

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 membersof the Apollo 13 Review activity.

2-34
Panel 1 conducted a formal interview with the MSCFlight Director

covering all significant mission events from the standpoint of ground

controllers.

Panels 2 through 4 continued developing preliminary reports. Panel 4

announceda formal schedule of interviews of MSC,contractors, and NASA

Headquarters personnel.

Board Membersexplored in detail possible failure modesequences

developed by MSCpersonnel involving ignition and combustion within the

SMcryogenic 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 further

activity for the week of May ll.

MAY9, 1970

Board in recess.

MAYi0, 1970

Board in recess.

MAYii, 1970

Board in recess. MSCsupport personnel continued work obtaining

additional technical data for Board review.

MAY12, 1970

Board Membersreturned to MSC.

Board Membersattended a General Session to review progress and

status of the report.

Panel Chairmenreported on individual progress of work and established schedules for completion of analyses and evaluations.

Chairman Cortright reported on the Langley Research Center support

test program aimed at simulation of SMpanel ejection energy pulses.

2-35
MAY13, 1970

Board Membersreviewed preliminary drafts of report chapter on Review and Analysis and Panel i report on Mission Events.

Mission Events Panel Membersinterviewed Electrical, Electronic, and

CommunicationsEngineer (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 ManagementPanel, conducted interviews with principal Apollo 13 program personnel from MSCand contract organizations.

Panel Memberscontinued drafting preliminary versions of Panel reports for review by the Board.

Manufacturing and Test Panel representatives discussed program for

oxygen tank testing to be conducted at BeechAircraft.

Board Membersmet in General Session to review report milestones and

required test data for the week ahead.

MAY14, 1970

Board met in General Session to review Panel report progress and to

agree to firm schedules for completion of all Review Board assignments.

Project ManagementPanel continued to interview key Apollo project

personnel from NASACenters and contractors.

Panel Memberscirculated first drafts of all Panel reports to Board

Membersfor review and correction.

MAY15, 1970

Mission Events Panel personnel interviewed Apollo 13 CommandModule

Pilot John Swigert to verify event chronology compiled by the Panel and

to review crew responses during Apollo 13 mission.

Project ManagementPanel continued interviewing key project personnel

with NASACenters and contractors.

2-36
.° MSC personnel provide Board Members and Panel Chairmen with a detailed briefing on all support tests and analyses being performed in

connection with the MSC and Board reviews.

Board Members met in Executive Session to review preliminary drafts

of 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 to

compile reference data in support of findings.

MAY 16, 1970

Board met in General Session to review further revisions of preliminary findings and determinations and to establish working schedules for

completion of 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 documentation.

MAY 17, 1970

Draft material for all parts of Board report was reviewed by Panel

Members 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 installations.

The Apollo 13 Review 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

review by Board Members.

Board reviewed a preliminary draft of findings and determinations

prepared by Panel Chairmen, Board Members,and Board Chairman.

A Manufacture and Test Panel representative reviewed special oxygen

tank test programs at BeechAircraft.

MAY19, 1970

Board Membersmet in Executive Session to continue evaluation and

assessment of preliminary findings, determinations, and recommendations

prepared by individual Board Membersand Panel Chairmen.

Board met in General Session to review final draft of Mission Events

Panel report.

Manufacture and Test Panel preliminary report was distributed to

Board Members for review and comment.

Design Panel preliminary report was distributed to Board Members for

review and comment.

Design Panel Members met with MSC Team officials to discuss further

test and analyses support for the Board.

MAY 20, 1970

Board Members met in Executive Session to review and evaluate reports

from the Design Panel and from the Manufacturing and Test Panel.

Project Management Panel distributed final draft of its report to

Board Members for review and comment.

Chairman Cortright met withMr. Bruce Lundin of the Aerospace Safety

Advisory Panel to discuss progress of Board review and analysis.

MAY 21, 1970

Board Members met in Executive Session for final review of Project

Management Panel report.

2-38
Board Membersand others met with MSCofficials to review in detail

the activities and actions taken after the Apollo 204 accident concerning

ignition flammability for materials and control in the CSM.

A third draft of preliminary findings, determinations_ and recommendations was developed and circulated by the Chairmanfor review and

comment.

Arrangements were madewith NASAHeadquarters officials for packaging, delivery, and distribution of the Board's final report.

Mission Events Panel conducted an interview with Lunar Module Pilot

Haise to review selected mission events bearing on the accident.

MAY22, 1970

Mission Events Panel representatives met with MSCofficials to review

in detail several events which occurred during later flight stages.

Board 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.

Langley Research Center representative M. Ellis briefed the Board on

ignition and combustion of materials in oxygen atmosphere tests being conducted in support of the Apollo 13 Review.

Board Observer I. I. Pinkel briefed the Board on Lewis Research

Center fire propagation tests involving Teflon.

MAY23, !970

Board Membersreviewed Chapter 4 of Board report entitled "Review

and Analysis."

Panel Chairmen reviewed draft findings and determinations prepared

by the Board.

2-39

N N L L K L.',". L . h 1_: L . _ I: _'" li /d< I:: L _ L
MAY 24, 1970

Board Members reviewed NASA Aerospace Safety Panel report covering

Apollo activities during the period of 1968-69.

Board met in Executive Session for detailed review of support test

status and progress and of documentation describing the results of test

activity.

Board met in Executive Sessicn for further review of findings,

determinations, and recommendations.

MAY 25, 1970

Board met in Executive Session to review test progress and decided

to postpone submittal of final report until June 8 in order to consider

results of Langley Research Center panel ejection tests.

Board Members continued to review MSC Investigation Team preliminary

drafts and refine Apollo 13 data in the various Board appendices.

Board met in Executive Session for further consideration of findings,

determinations, and recommendations.

MAY 26, 1970

Board met in General Session and interviewed Astronaut James Lovell

regarding crew understanding of inflight accident.

Board Members reviewed proposed MSC tank combustion test and agreed

to test methodology and objectives.

Panel Members continued preparation of individual Panel reports.

MAY 27, 1970

Board and Panel Members received a detailed briefing on thermostatic

switch failure during MSC 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
Board met in General Session to review status of Panel reports,

documentation of test data and results, and plans for report typing and

review.

Board agreed to recess for several days to accumulate additional

test information on panel separation and full scale tank ignition data.

MAY28, 1970

Board in recess.

_Y 29, 1970

Board in recess.

MAY30, 1970

Board in recess.

MAY31, 1970

Board in recess.

JUNEi, 1970

Board Membersreturned to MSC.

Board and Panel Membersmet in General Session to discuss revisions

of Panel reports in light of latest information regarding thermostatic

switch failure during CDDTat KSC.

Board approved new schedule for Board report calling for final

versions of Panel reports by Monday, June 8.

2-41
JUNE2, 1970

Chairman Cortright briefed the Press on the status of the Board's

work and future plans.

Board and Panel Membersparticipated in a detailed interview and

discussion with MSCand contractor personnel regarding specific coordination steps taken during oxygen tank no. 2 detanking operations at KSC.

Board Membersmet in Executive Session to review latest test results

and to assess status of Board findings and determinations.

JUNE3, 1970

Board and Panel Membersmet with MSCProgram Office personnel for a

detailed update of recent MSCinformation and analyses stemming from ongoing test programs.

Board Membersand Panel Chairmen completed final reviews of Panel

reports and also reviewed final draft of findings, determinations, and

recommendations.

Board and Panel Membersreceived a detailed briefing on thermostatic

switch questions with emphasis upon actions of various organizations

during and after detanking operations at KSC.

JUNE4, 1970

Board Membersmet in Executive Session and completed final revisions

of Chapter 4 of the Board summary.

Board and Panel Memberswitnessed 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
- JUNE5, 1970

Board Members met in Executive Session and completed work on Chapter 5 of the Board Summary Report (Findings, Determinations, and Recommendations).

Board Members reviewed final version of Project Management Panel

report and authorized printing as Appendix E.

Board Members Hedrick and Mark completed final tabulation of test

support activities performed for the Board.

Board Members reviewed films of special test activities performed

at various NASA Centers.

JUNE 6, 1970

Board met in Executive Session throughout the day and completed

its review of Chapter 5 of its report (Findings, Determinations, and

Recommendations).

Board Members completed review of analyses to be incorporated in

Appendix F, Special Tests and Analyses.

JUNE 7, 1970

The Board met in Executive Session and approved plans and schedules

for final editorial review and publication of the Board report.

The Chairman recessed the Board until June 15 at which time the

Board is scheduled to reconvene in Washington, D.C., to present its

report to the NASA Administrator and Deputy Administrator.

2-43
This page left blank intentionally.

2-44
CHAPTER 3

DESCRIPTION OF APOLLO 13 SPACE VEHICLE

AND MISSION SUMMARY

3-0

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 functions and capabilities.

Part 2 describes the Apollo 13 mission and gives a mission sequence

of events summary.

3-1
PART i APOLL0/SATURN V SPACE VEHICLE

The primary flight hardware of the Apollo Program consists of the

Saturn V launch vehicle and Apollo spacecraft (fig. 3-1). Collectively,

they are designated the Apollo/Saturn V space vehicle (SV). Selected

major systems and subsystems of the space vehicle may be summarized as

follows.

SATURN 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-II, 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-I

rocket engines. These engines develop a nominal sea level thrust total

of approximately 7,650,000 pounds. The stage dry weight is approximately

288,000 pounds and the total loaded stage weight is approximately

5,031,500 pounds. The S-IC stage interfaces structurally and electrically with the S-II stage. It also interfaces structurally, electrically, and pneumatically with ground support equipment (GSE) through

two umbilical service arms, three tail service masts, and certain

electronic systems by antennas. The S-IC stage is instrumented for

operational measurements or signals which are transmitted by its independent telemetry system.

S-II Stage

The S-II 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-II stage is

approximately 78,050 pounds. The stage approximate loaded gross weight

is 1,075,000 pounds. The S-IC/S-II interstage weighs 10,460 pounds.

The S-II stage is instrumented for operational and research and development measurements which are transmitted by its independent telemetry

system. The S-II stage has structural and electrical interfaces with

the S-IC and S-IVB stages, and electric, pneumatic, and fluid interfaces

with GSE through its umbilicals and antennas.

3-2
rr::.

r:.

rL_

I

L_J

aunch escape system

/_._Boost protective cover

Commandmodule

iiiiiiiill

Lunar module

adapter

Spacecraft

Figure

32 ft

1

363 ft

Space vehicle

5-i.- Apollo/Saturn V space vehicle.

Instrument unit

S-IVB

Inter-stage

S-II

Inter-stage

S-IC

LAUNCHVEHICLE

Launch vehicle

..,.

..
FLIGHT TERMINATION

RECEIVERS (2)

GOX

DISTRIBUTOR

INSTRUMENTATI ON

HELIUM

CYLINDERS (4)

LINE

FT

.7 IN

FORWARD

SKIRT

IN OXIDIZER

TANK

ANNULAR

BAFFLES

CENTER

ENGINE

SUPP(

FUEL

SUCTION

HEAT

LOWER

THRUST RIN

F-l ENGINES

(5)

INSTRUMENTATIC

RETROROCKETS

BAFFLE

LINE

TUNNELS (5)

SUCTION

LINES (5)

TUNNEL

THRUST

RI NG

HEA

FLIGHT CONTROL

SERVOACTUATOR

262.4 IN

INTERTANK

SECTION

FUEL

IN TANK

Figure 3-2.- S-IC stage.

3-4
VEHI CLE

STATION

2519

81-I/2

FEET

;YSTEMSTUNNEL

T FORWARDSKI RT

11-I/2 FEET

LIQUID HYDROGEN

TANK

(37,737 CU FT)

;EET

LH2/LOX COMMON

BULKHEAD

LIQUID OXYGEN

(12,745.5 CU FT)

22 FEET

VEHICLE

STATION

1541 33 FEET _ I

AFT SKIRT

,____1 THRUST STRUCTURE

14-I/2 FEET

_[ INTERSTAGE

18-I/4 FEET

Figure 3-3.- S-II stage.

3-5
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 J-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 launch weight of the stage is 261,700 pounds.

The interstage weight of 8100 pounds is not included in the stated

weights. The stage is instrumented for functional measurements or signals which are transmitted by its independent telemetry system.

The high performance J-2 engine as installed in the S-IVB stage

has a multiple start capability. The S-IVB J-2 engine is scheduled

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 4.5:1)

during the first i00 seconds of translunar injection. The remaining

translunar injection acceleration is provided at a thrust level of

203,000 pounds (mixture mass ratio of 5.0:1). 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.

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.

APOLLO 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. The

Apollo S/C consists of the spacecraft-to-LM adapter (SLA), the service

module (SM), the command module (CM), the launch escape system (LES), and

the lunar module (LM). The CM and SM as a unit are referred to as the

command and service module (CSM).

3-6
59.0

FEET

i.-..--21.6 FEET-I_

FORWARD SKIRT

LH2 TANK

I0,418

CU FT--

mlI0.2 FEET

L

44.0 FEET

LOX TANK

2830 k/ \_J_

AFT SKIRT _"/!. _ " ]

)

THRUST STRUCTURE _'

(WITH ENGINE 5.2tFEET

ATTACHED)

, 33.0 FEET =I

AFT INTERSTAGE

19 FEET

Figure 3-4.- S-IVB stage.

3-7

ILl N _ E l/ E E L L 11. L.' L: U II g l: n L _.
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. Aerodynamically, 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.

Spring thrusters are used to separate the LM 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. Simultaneously, four

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.

Service Module

The service module (SM)(fig. 3-6) provides the main spacecraft propulsion and maneuvering capability during a mission. The SM provides

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.

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. Three 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 gap between

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.

3-8

ILl ILl 13 L L 1J ]2 L " _ L_ L _ _-_
I

UPPER (FORWARD)

21 ' JETTISONABLE

PANELS

(4 PLACES)

LOWER (AFT)

7' FIXED PANELS

CIRCUMFERENTIAL

LINEAR-SHAPED CHARGE

LONGITUDINAL

LINEAR-SHAPED CHARGE

(4 PLACES)

PYROTECHN IC THRUSTERS

(4 PLACES)

CIRCUMFERENTIAL

LINEAR-SHAPED CHARGE

THRUSTER/HINGE

(2) (4 PLACES)

IU

Figure 3-5.- Spacecraft-to-LM adapter.

3-9
RED ELECTRICAL

DOCKING POWER

LIGHT SUBSYSTEM

RADIATORS

SM REACTION

CONTROL FLYAWAY

SUBSYSTEM UMBILICAL

QUAD

FLOODLIGHT

GREEN

SCIMITAR DOCKING

ANT LIGHT

ENVIRONMENTAL

CONTR

RADIATOR

EXTENSION

FUEL TANKS

\

FORWARD BULKHEAD INSTALL,

FUEL CELLS

PRESSURIZATION (_ [_

OXYGEN TANKS

HYDROGEN TANKS _

S-BAND HIGH GAIN ANTENNA

12 FT 10 IN. . i SERVICE PROPULSION ENGINE

SECTOR 2 _ SERVICE PROPULSION SUBSYSTEM

SECTOR 3 _ OXIDIZER TANKS

SECTOR 4 OXYGEN TANKS, HYDROGEN TANKS, FUEL CELLS

SECTOR 5 _ SERVICE PROPULSION SUBSYSTEM

SECTOR6 J FUEL TANKS

CENTER SECTION- SERVICE PROPULSION ENGINE AND

HELIUM TANKS

UM TANKS

__RE'A OXIDIZER

TANKS

CTION

_1 ICONTROL

J-_ ISUBSYSTEM

I _QUAOS 141

AFT

BULKHEAD

Figure 3-6.- Service module.

3-10

ILl ILl [ L L L: LL L L I_ la L' _' u E E -
. -

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 radiator 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.

Propulsion.- Main spacecraft propulsion is provided by the

20500-pound thrust SPS. The SPS engine is a restartable, non-throttleable

engine which uses nitrogen tetroxide (N204) as an oxidizer and a 50-50

mixture of hydrazine and unsymmetrical-dimethylhydrazine (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 assembly 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. The SM RCS provides for maneuvering

about and along three axes.

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 _ 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
Seven lights are mounted in the aluminumpanels 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). The inner structure, the pressurized crew compartment, is made

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.

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 LM

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 controls 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
+X ÷y

_y ,_ __< --Z

/ COMBINED TUNNEL HATCH

_ AUNCH ESCAPE TOWER ATTACHMENT (TYPICAL)

_ _"._'_'E7 P,TC.

(, ATTACHPOINT

(TYPICAL)

+X +y

-Y -X -Z

LEFT HAND

FORWARD COMPARTMENT FORWARD EQUIPMENT BAY COMBINED TUNNEL HATCH

CREW ".OWER , FORWARD " _ _ _ _OMPARTMENT

{TYPICAL)

HAND EQUIPMENT BAY RIGHT HAND EQUIPMENT /

AFT COMPARTMENT AFT COMPARTMENT

Figure 3-7.- Command module.

3-13
Audio J

control--_//_

contro/-_7 / Flightl

_f_'_--SCS pane" ' power I

Cry°genics- 7 /--Service

/ / propulsion

r a C autai:nln-g-r_-__.> _ /.--Audio

F,'I coo''o'

RCSmgmt ECS I I........._ i \

I c0_t,r__L_l I=';owre_'"k________

\ I '

/ _E n vC_i;::ie nla i cJontr o i__

• Launch vehicle emergency detection

• Flight attitude

• Mission sequence

• Velocity change monitor

• Entry monitor

• Propellant gauging

• Environment control

• Communications control

• Power distribution

• Caution and warning

LIJJ "--J'_-]-I I _ --':--:

I Flight controls p. Systems controls

I --_l=-_,LII ,_--_ll . "'_JJaJJl_l ul'IrlI_

/-J71 ', ," ', ," I\ ..X

• J • •

% • % •

% • • ••

•S

_o sx

• % S %

I •

Commander CM pilot LM pilot

Figure 3-8.- CM main display console,

3-14
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 abnormality. 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 (LM) (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

descent stage. The stages are attached at four fittings by explosive

bolts. Separable umbilicals and hardline connections provide subsystem

continuity to operate both stages as a single unit until separate ascent

stage operation is desired. The LM is designed to operate for 48 hours

after separation from the CSM, with a maximum lunar stay time of 44 hours.

Table 3-I is a weight summary of the Apollo/Saturn 5 space vehicle for

the Apollo 13 mission.

3-15
S-band

steerable

an

Rendezvous

radar antenna_

S-band in-flight

antenna (2)

Docking

light (

Overhead

hatch

Docking Ascent VHF

window stage antenna

EVA

antenna

Docking

target

equipment

bay

RCS thrust chamber

assembly cluster

Tracking I

Forward hatch

Forward,

Docking

ight (3)

Landing gear

Ladder Egress Descent Descent Landing

platform engine stage radar_

skirt antenna

Figure 3-9.- Lunar module.

3-]_6
TABLE 3-I.- APOLLO 13 WEIGHT SUMMARY (WEIGHT IN POUNDS)

Stage/module

S-IC

s-lc/s-ii

interstage

S-If stage

S-II/S-IVB

interstage

S-IVB stage

Instrument unit

Inert weight

288OOO

11464

78050

Total

expendables

4746870

996960

Total weight

5034870

11464

1075010

8100

25050

4482

236671

8100

261721

4482

Final

separation

weight

363403

92523

35526

Launch vehicle at ignition 6,395,647

Spacecraft-LM

adapter

Lunar module

Service module

Command module

Launch escape

system

4o44

9915

10532

12572

9012

23568

40567

4044

33483

51099

12572

9012

"33941

**14076

**11269

(Landing)

* CSM/LM separation

** CM/SM separation

3-17
" TABLE 3-1.- APOLLO 13 WEIGHT SUMMARY (WEIGHT IN POUNDS) - Concluded

Stage/module Inert weight

Total

expendables

Final

Total weight separation

weight

Spacecraft at ignition 110,210

Space vehicle at ignition

S-IC thrust buildup

Space vehicle at lift-off

Space vehicle at orbit insertion

6505857

(-)84598

6421259

299998

Main propulsion.- Main propulsion is provided by the descent propulsion system (DPS) and the 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. The descent

engine is gimbaled, throttleable, and restartable. The engine can be

throttled from 1050 pounds of thrust to 6300 pounds. Throttle positions

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.

3-18
The AGScommandson/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, commandmore thrust than the level commandedby the LMguidance computer

(LGC).

The ascent engine is a fixed, non-throttleable engine. The engine

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.

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). During the lunar descent phase,

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. All batteries and busses may be individually monitored for

load, voltage, and failure. Several isolation and combination modes

are provided.

Two inverters, each capable of supplying full load, convert the

dc to ac for ll5-volt, 400-hertz supply. Electrical power is distributed

by the following busses: LM Pilot's dc bus, Commander's dc bus, and ac

busses A and B.

The four descent stage silver-zinc batteries are identical and have

a 400 ampere-hour capacity at 28 volts. Because the batteries do not

have a constant voltage at various states of charge/load levels, "high"

and "low" voltage taps are provided for selection. The "low voltage"

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

3-19
_apacity 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

componentsdesigned into the EPSto enhanceEPSreliability. 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 maximumdrain of 50 ampsper battery and descent energy is

46.9 kilowatt hours at a maximumdrain of 25 ampsper battery.

MISSIONMONITORING,SUPPORT,ANDCONTROL

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; commandingthe space vehicle by transmitting

real-time and updata commandsto 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 sametime that the data

event occurred.

The flight crew and the following organizations and facilities

participate in mission control operations:

a. Mission Control Center (MCC),MannedSpacecraft Center (MSC),

Houston, Texas. The MCCcontains the communication, computer display,

and commandsystems to enable the flight controllers to effectively

monitor and control the space vehicle.

b. KennedySpace Center (KSC), CapeKennedy, Florida. The space

vehicle is launched from KSCand controlled from the Launch Control

Center (LCC). Prelaunch, launch, and powered flight data are collected

at the Central Instrumentation Facility (CIF) at KSCfrom the launch

pads, CIF receivers, Merritt Island LaunchArea (MILA), and the downrange Air Force Eastern Test Range (AFETR)stations. These data are

3-20
transmitted to MCCvia the Apollo LaunchData"System (ALDS). Also

located at KSC(AFETR)is the Impact Predictor (IP), for range safety

purposes.

c. GoddardSpaceFlight Center (GSFC), Greenbelt, Maryland. GSFC

managesand operates the MannedSpace Flight Network (MSFN)and the

NASAcommunications (NASCOM)network. During flight, the MSFNis

under the operational control of the MCC.

d. George C. Marshall SpaceFlight Center (MSFC),Huntsville,

Alabama. MSFC,by meansof the Launch Information ExchangeFacility

(LIEF) and the Huntsville Operations Support Center (HOSC)provides

launch vehicle systems real-time support to KSCand MCCfor 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 MCChas 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 mannedby flight control specialists who provide detailed support to the MOCR. Figure 3-11 outlines

the organization of the MCCfor flight control and briefly describes

key responsibilities. Information flow within the MOCRis shownin

figure 3-12.

The consoles within the MOCRand 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 recommendor 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
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,
rr=

r_

r:

O0

I

h)

tO

Goddard Houston LIEF Marshall

ALDS

Kennedy AFETR

ALDS - Apollo Launch Data System

LIEF - Launch Information Exchange Facility

.m,

Figure 3-i0.- Basic telemetry, command, and communication

interfaces for flight control.
PUBLIC AFFAIRS

MISSION STATUS

TO PUBLIC

MISSION COMMAND

AND CONTROL GROUP

OPERATIONS & PROCEDURES (o&P)I

MCCIMSFN MISSION CON- I

TROL PROCEDURES; FLIGHT L_

CONTROL SCHE DU LING; MANNING;

CONTROL FORMAT; DISPLAYS; |

TELETYPE TRAFFIC ANALYSIS |

NETWORK £ONTROL

MSFN CONTROL, RADAR AND

COMMAND HAN DOVERS

SPACECRAFT COMMUNICATOR

COMMUNICATIONS

WITH SPACECRAFT

FLIGHT ACTIVITIES (FAD)

FLIGHT PLAN DETAILED

IMPLEMENTATION

I SPACE ENVIRONMENT (SEO)

SPACE RADIATION

ENVIRONMENT DATA

I XPERIMENT ACTIVITIES MEAD)

INF LIGHT EXPERIMENT

IMPLEMENTATION

I

I I

DIRECTOR SSR

SSR

MISSION DIRECTOR )MD) ]

I

OVERALL CONDUCT OF

MISSION

I

I FLIGHT DIRECTOR (FD)

DECISIONS/ACTIONS ON SPACE

VEHICLE SYSTE MS/DYNAMICS

AND MCC/MSFN OPERATIONS

I

SYSTEMS OPERATIONS

GROUP

._BOOSTER SYSTEMS ENGINEERS (BSE_

MONITOR STATUS OF L._

S.IC, S-II, S-IVB FLIGHT I-_

SYSTEMS /

I

I HUNTSVILLE OPERATIONS I

SUPPORT CENTER (HOSC) I

._ EMU ENGINEERS

COMPUTE R UPDATE OF

CONSUMABLES DATA;

EVA DECISIONS

I

LIFE 1

SYSTEMS

SSR

SPACECRAFT SYSTEMS ENGINEERS

MONITOR STATUS OF

E LECTRICAL. COMMUNICATION,

INSTRUMENTATION. SEQUENTIAL,

LIFE SUPPORT. STABILIZATION

AND CONTROL. PROPULSION. AND

GUIDANCE AND NAVIGATION

SYS'I_ MS

LIFE SYSTEMS (SURGEON)

MONITORS PHYSIOLOGICAL AND

ENVIRONMENTAL STATUS OF

FLIGHT CREW

I

PROGRAM

OFFICE

I SC PLANNING I

AND ANALYSIS

SSR

I

i.=,o, i

EVALUATION

ROOM

I OD MANAGER I

RECOVERY AND OTHER

MISSION SUPPORT

FLIGHT DYNAMICS

GROUP

FLIGHT DYNAMICS OFFICER (FDO)

iMONITORS PRE LAUNCH CHECKOUT

POWERED FLIGHT EVENTS AND

TRAJECTORIES; REENTRY EVENTS

AND TRAJECTORIES

RETROFIRE OFFICER (RETRO)

MAINTAINS UPDATED ABORT

AND REENTRY PLAN; UPDATES

IMPACT POINT ESTIMATES

t GUIDANCE OFFICER (GUIDO)

MONITORS GUIDANCE

FUNCTIONS DURING POWERED

FLIGHT AND PREMANEUVER

PREPARATION

I

I II J

VEHICLE FLIGHT

SYSTEMS DYNAMICS

SSR SSR

KSCLAUNCH

OPERATIONS

Figure B-ll.- Mission Control Center organization.

3-2)4
LAUNCH

VEHICLE

STAGES

STAGE STATUS

VEHI CLE

SYSTEMS

I

I

SYSTEMSSTATUS

ASSISTANT

FLIGHT

DIRECTOR

MCC/MSFN

STATUS

MISSION

PROCEDURE

STATUS

MISSION PROCEDURESTATUS

I 0OFFICER AND P I

u")

MISSION

DI RECTOR

_L

f,Z)

LU 0

I:C _-w

r'_ V')

FLIGHT DIRECTOR

EQUIPMENT STATUS

FLIGHT

INFORMATION

II: STATUS NETWORK

I SlC DATA

L FLIGHT i_

CREW J

M AND 0

SUPERVISOR

FLIGHT

DYNAMICS

GROUP

NETWORK

CONTROLLER

i SPACECRAFT COMMUNICATOR

rl,

Figure 3-12.- Information flc_within the

Mission Operations Control Room.

3-25
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 Mauro Formation.

Deploy and activate an Apollo Lunar Surface Experiments Package

(m_SEP).

Develop man's capability to work in the lunar environment.

Obtain photographs of candidate exploration sites.

Table 3-II lists the Apollo 13 mission sequence of major events and

the time of occurrence in ground elapsed time.

TABLE 3-II. - APOLLO 13 MISSION SEQUENCE OF EVENTS

Event

Range zero (02:13:00.0 p.m.e.s.t., April ll)

Earth parking orbit insertion

Second S-IVB ignition

Translunar 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

Landing

Ground elapsed time

(hr:min:sec)

00:00:00

00:12:40

02:35:46

02:41:47

03:06:39

04:01:03

04:18:01

05:59:59

30:40:50

55:54:53

61:29:43

77:56:40

79:27:39

105:18:32

137:39:49

138:02:06

141:30:02

142:40:47

142:54:41

3-26
Launch and Earth Parking Orbit

Apollo 1B was successfully launched on schedule from Launch Complex

39A, KennedySpace Center, Florida, at 2:13 p.m.e.s.t., April ll, 1970.

The launch vehicle stages inserted the S-IVB/instrument unit (IU)/

spacecraft combination into an earth parking orbit with an apogeeof

100.2 nautical miles (n. mi. ) and a perigee of 98.0 n. mi. (100-n. mi.

circular planned). During second stage boost, the center engine of the

S-II stage cut off about 132 seconds early, causing the remaining four

engines to burn approximately 3_ seconds longer than predicted. Space

vehicle velocity after S-II boost was 223 feet per second (fps) lower

than planned. As a result, the S-IVB orbital insertion burn was approximately 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. A greater than B-sigma probability of

meeting translunar injection (TLI) cutoff conditions existed with remaining S-IVB propellants.

After orbital insertion, all launch vehicle and spacecraft systems

were verified and preparation was madefor translunar injection (TLI).

Onboardtelevision was initiated at 01:35 ground elapsed time (g.e.t.)

for about 5.5 minutes. The second S-IVBburn was initiated on schedule

for TLI. All major systems operated satisfactorily and all end conditions were nominal for a free-return circumlunar trajectory.

Translunar Coast

The CSMseparated from the LM/IU/S-IVB at about 03:07 g.e.t. Onboard television was then initiated for about 72 minutes and clearly

showedCSM"hard docking," ejection of the CSM/LMfrom the S-IVB at

about 0_:01 g.e.t., and the S-IVB auxiliary propulsion system (APS)

evasive maneuveras well as spacecraft interior and exterior scenes.

The SMRCSpropellant usage for the separation, transposition, docking,

and ejection was nominal. All launch vehicle safing activities were

performed as scheduled.

The S-IVBAPS evasive maneuverby an 8-secondAPS Ullage burn was

initiated at 04:18 g.e.t, and was successfully completed. The liquid

oxygen dumpwas initiated at 04:39 g.e.t, and was also successfully

accomplished. The first S-IVB APSburn for lunar target point impact

was initiated at 06:00 g.e.t. The burn duration was 217 seconds, producing a differential velocity of approximately 28 fps. Tracking information available at 08:00 g.e.t, indicated that the S-IVB/IU would impact

at 6°53' S., B0°5B' W. versus the targeted B° S., B0° W. Therefore, the

second S-IVB APS(trim) burn was not required. The gaseousnitrogen pressure dropped in the IU ST-124-MBinertial platform at 18:25 g.e.t, and

the S-IVB/IU no longer had attitude control but began tumbling slowly.

3-27
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. The S-IVB/IU impacted the

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. The targeted impact point was

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 accuracy of the TLI maneuver was such that spacecraft midcourse

correction No. 1 (MCC-1), scheduled for ll:41 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. All SPS burn

parameters were normal. The accuracy of MCC-3 was such that MCC-3,

scheduled for 55:26 g.e.t., was not performed. Good quality television

coverage of the preparations and performance of MCC-2 was received for

49 minutes beginning at 30:13 g.e.t.

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 152:00 g.e.t.

Transearth Coast

At pericynthion plus 2 hours (79:28 g.e.t.), a LM DPS maneuver was

performed to shorten the return trip time and move the earth landing

point. The 263.4-second burn produced a differential velocity of 860.5

fps and resulted in an initial predicted earth landing point in the midPacific Ocean at 1_2:53 g.e.t. Both LM guidance systems were powered

up and the primary system was used for this maneuver. Following the

maneuver, passive thermal control was established and the LMwas powered

down to conserve consumables; only the LM environmental control system

(ECS) and communications and telemetry systems were kept powered up.

The LMDPS 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 crew viewed and photographed

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.

Entry and Recovery

Weather in the prime recovery area was as follows: broken stratus

clouds at 2000 feet; visibility l0 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 JIMA. The crew, picked up by a recovery helicopter, was safe aboard the ship at 1:53 p.m.e.s.t., less than an hour

after landing.

3-29

U 1/- IL L: E. L L E L. ta U n n u L
I

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3-30
CHAPTER 4

REVIEW AND ANALYSIS OF APOLLO 15 ACCIDENT

4-0
ILl li It"., /_ IL 1: 12 L ' _: I_ L'_ _' ,_ 13 E IL _ L.. _
PART i. 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,

ll, 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 II 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 drawing 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 tank contents.

The other, called the quantity probe, consists of an upper section which

supports a cylindrical capacitance 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 wiring 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 not reach from the fill line to the gage cylinder (fig. 4-3). Thus, the variations might be such that a very loose

fit would result.

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 command module (CM). The supply line also connects

4-2
• °

Blowout disc

Closeout cap

Supply

line

Temperature

sensor

Fan

motor

"hermostat

Heater

To fuel

ceil/ECS'_

Capacitance

gage

Insulation

Pressure

transducer

Pressure

switch

Relief

valve

Overboard

Fan

motor

Closeout cap

Figure &-l.- Oxygen tank no. 2 internal components.

tl !1 L .... ._ __ __
Fill tube conduit

Electrical

wiring

Supply

Oxygen

vent tube

Filter

O0

O0

Temperature sensor _ /i

fan 1

Quantity probe

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

4-4
r

I

e

• Dimension a

depends on

value of e

Part Max Nom Min

dim dim dim

a :_ 0.28 0.24 0.16

b 1.095!1.0801.065

c 0.26 0.20 0.14

d 1.45 1.43 1.41

e 18 ° 21 ° 24 °

Adverse

tolerance case

Fill

tube

I

\

\ I

\

\rI

Probe

Nominal

tolerance case

Fill

tube

\

r-:

I"

Figure 4-3.- Nominal and adverse tolerance cases.
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 -297 ° F and operates over the range from -340 ° F to +80 ° F.

The term "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 oxygen 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 i000 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.

TABLE 4-1.- MATERIALS WITHIN OXYGEN TANK

Mat eri al

Teflon-wire insulation

sleeving and solid

Aluminum (all forms )

Stainless steel

Inconel alloys

Approximate

quantity, lb

1.1

0.8

2.4

1.7

Available

energy, Btu

2,400

20,500

15,000

2,900

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

• _
f

_Fuel cell

I

Oxygen I l

tank 2"-_

Oxygen subsystem

shelf module

/drogen subs

_.she If module

Fuel

cell

3

:uel

cell

2

Fuel

cell

1

-Oxygen

tank 1

- Oxygen valve

module

Oxygen servicing

panel

.Hydrogen

1

tank 2

Fisure_-4. - Arrangement of fuel cells and cryogenic systems in bay h.

4-7

I1 N E :" E :I2 L L K I:: _. " I: g. 1i _ t: iz L ± ,__
This page left blank intentionally.

_-8

ILl ILl 13 L L ld 12 k L _ '_ '_ "
/_.S _..... ., _._<.c

/ F t J -_

Figure 4=5.= Fuel cells shelf.

4-9

.,¢- _* _<-C - L -_C,2 / L, '

Figure 4-6.- Oxygen tank shelf.

Precedingpageblank

4-11

L k E '_ k 'L. '_ L" L" L_ /Li _ K L

FigureX_=_XX Hydrogen tank shelf.

Precedingpageblank

4-13

ILl li E " L: E L L
. i_

E l.:: '
F

4_8° --

Figure _X_ Inside view of panel covering bay 4.

Precedingpageblank

4-15

of the Apollo 13 service module (SM109) at the North American Rockwell

plant prior to shipment to KSC. Figure 4-5 showsthe fuel cell shelf,

with fuel cell 1 on the right, fuel cell 3 on the left, and fuel cell 2

behind cells 1 and 3. The top of oxygen tank no. 2 can be seen at the

lower left. Figure 4-6 showsthe oxygen tank shelf, with oxygen tank

no. 2 at left center. Figure h-7 shows the hydrogen tank shelf with

hydrogen tank no. 1 on top and hydrogen tank no. 2 below. The bottom

of the oxygen shelf shows someof the oxygen system instrumentation and

wiring, largely covered by insulation. Figure h-8 is a photograph of

the bay 4 panel, which was missing from the service module after the

accident.

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 madeby SimmondsPrecision 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 10024XTA0008. It was the eighth Block II oxygen

tank built. Twenty-eight Block I oxygen tanks had previously been built

by Beech.

The design of the oxygen tank is such that once the upper and lower

halves of the inner and outer shells are assembledand welded, the

heater assembly must be inserted in the tank, movedto 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 movementinside the tank, where

movementcannot be readily observed, and where possible damageto 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. Rewelding was

also required when it was determined that incorrect welding wire had

been inadvertently used for a small weld on a vacuumpumpmounted on

Precedingpageblank

4-17
the outside of the tank dome. The upper fan motor originally installed

_as 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 pumpeddown

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.

After 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

R00 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. h-2). No difficulties were recorded in this

det anking operation.

The acceptance test indicated that the rate of heat leak into the

taukwas higher than permitted by the specifications. After some reworking, the rate improved, but was still somewhat higher than specified.

The tank was accepted with a formal waiver of this condition. 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 May 3, 1967.

ASSEMBLY AND TEST AT NORTH AMERICAN ROCKWELL

The assembly of oxygen shelf serial number 0632AAG3277, with Beech

oxygen tank serial number 10024XTA0009 as oxygen tank no. 1 and serial

number 1002_XTA0008 as oxygen tank no. 2, was completed on March ll, 1968.

The shelf was to be installed in SM 106 for flight in the Apollo l0

mission.

4-18
Beginning on April 27, the assembled oxygen shelf underwent standard proof-pressure, leak, and functional checks. Onevalve 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 tauk no. 2 was

required. None of the oxygen tank testing at NR requires use of L0X

in the tanks.

On June _, 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. 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.

The oxygen shelf was removed in the manner shown 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 h. 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 106 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_ithout further difficulty. Following

removal, the oxygen shelf was retested to check shelf integrity, including proof-pressure tests, leak tests, and functional tests of

pressure transducers and switches, thermal switches, and vac-ion pumps.

No cryogenic testing was conducted. 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 no. 2. Further calculations and tests conducted during this investigation, however, have indicated that the

forces experienced by the shelf were probably close to those originally

4-19
rr_

|

ro

o

--- Fuel cell shelf

A

hoist

E

Sling

Sling adapter _"

stment of ballance weights

Figure 4-9.- Hoist and sling arrangement - oxygen shelf.
calculated 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 could have

been displaced by the event.

The shelf passed these tests and was installed in SM 109 on

November 22, 1968. The shelf tests accomplished earlier in SM 106

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.

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.

The 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. The oxygen tanks were

evacuated to 5mmHg 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.

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. As a first

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. Further discussions between KSC, MSC, NR,

and Beech personnel considered thatthe 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. Referring to figure _-2, it will be noted

that such a leak would allow the gaseous oxygen (G0X) being supplied

to the vent line to leak directly to the fill line without forcing any

4-21
significant amount of LOXout 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 no. 1 emptied in a few minutes. Tank no. 2 did not. Additional

attempts were madewith higher pressures without effect, and a decision

was madeto try to "boil off" the remaining oxygen in tank no. 2 by

use of the tank heaters. The heaters were energized with the 65 V tic.

GSEpower supply, and, about l-l/2 hours later, the fans were turned

on to add more heat and mixing. After 6 hours of heater operation,

the quantity had only decreased to 35 percent, and it was decided to

attempt a pressure cycling technique. With the heaters and fans still

energized, the tank 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 about 8 hours of heater operation.

Suspecting the loosely fitting fill line connection to the quantity

probe inner cylinder, KSCpersonnel consulted with cognizant personnel

at MSCand at NRand 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 CMwould have been difficult

and would have taken at least 45 hours. In addition, shelf replacement

would have had the potential of damagingor degrading other elements of

the SMin the course of replacement activity. Therefore, the decision

was madeto test the ability to fill oxygen tank no. 2 on March 30,

1970, twelve days prior to the scheduled Saturday, April ll, launch,

so as to be in a position to decide on shelf replacement well before

the launch date.

Accordingly, flow tests with G0Xwere run on oxygen tank no. 2

and on oxygen tank no. 1 for comparison. No problems were encountered,

and the flow rates in the two tanks were similar. In addition, Beech

was asked to test the electrical energy level reached in the event of

a short circuit between plates of the quantity probe capacitance gage.

This test showedthat very low energy levels would result. On the

filling test, oxygen tanks no. 1 and no. 2 were filled with LOXto

about 20 percent of capacity on March 30 with no difficulty. Tank no. 1

emptied in the normal manner, but emptying oxygen tank no. 2 again

required pressure cycling with the heaters turned on.

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 NRwas not considered

and it was felt that the apparently normal detanking which had occurred

in 1967 at Beechwas not pertinent because it was believed that a

different procedure was used by Beech. In fact, however, the last

portion of the procedure was quite similar, although a slightly lower

G0Xpressure was utilized.

Throughout these considerations, which involved technical and

managementpersonnel of KSC,MSC,NR, Beech, and NASAHeadquarters,

emphasiswas directed toward the possibility and consequencesof 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 detanking sequences.

Manyof the principals in the discussions were not aware of the

extended heater operations. Thosethat did know the details of the

procedure did not consider the possibility of damagedue to excessive

heat within the tank, and therefore did not advise managementofficials

of any possible consequencesof the unusually long heater operations.

As noted earlier in this chapter, and shownin 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 MSCsince the accident,

however, it was found that the switches failed to openwhen the

heaters were powered from a 65 V dc supply similar to the power used

at KSCduring the detanking sequence. Subsequentinvestigations have

shownthat the thermostatic switches used, while rated as satisfactory

for the 28 V dc spacecraft power supply, could not openproperly at

65 V tic. 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 KSCindicates that, in fact,

the switches did not open when the temperature indication from within

the tank rose past 80° F. Further tests have shownthat the temperatures on the heater tube mayhave reached as muchas 1000° F during

the detanking. This temperature will cause serious damageto adjacent

Teflon insulation, and such damagealmost certainly occurred.

Noneof the above, however, Wasknown at the time and, after

extensive consideration was given to all possibilities of damagefrom

a loose fill tube, it was decided to leave the oxygen shelf and oxygen

tank no. 2 in the SMand 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
PART 3. THE APOLLO 13 FLIGHT

The Apollo 13 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. Candidate future landing sites were

scheduled to be photographed from lunar orbit with a high-resolution

topographic camera carried aboard the command module.

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. Lovell

and prime Lunar Module Pilot Fred Haise were immune to rubella, but that

prime Command Module Pilot Thomas K. Mattingly III 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 ll, 1970, from the

KSC Launch Complex 39A. The spacecraft was inserted into a lO0-nauticalmile circular earth orbit. The only significant launch phase anomaly was

premature shutdown of the center engine of the S-II second stage. As a

result, the remaining four S-II engines burned 34 seconds longer than

planned and the S-IVB third stage burned a few seconds longer than planned. At orbital insertion, the velocity was within 1.2 feet per second

of the planned velocity. Moreover, an adequate propellant margin was

maintained in the S-IVB for the translunar injection burn.

4-25
Orbital insertion was at 00:12:39 ground elapsed time (g.e.t.).

The 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 LaunchArea,

Florida, tracking station.

The S-IVB restarted at 02:35:46 g.e.t, for the translunar injection

burn, with shutdown coming some5 minutes 51 seconds later. Accuracy of

the Saturn V instrument unit guidance for the TLI burn was such that a

planned midcourse correction maneuverat 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.

The CSMwas separated from the S-IVB about 3 hours after launch,

and after a brief period of stationkeeping, the crew maneuveredthe CSM

to dock with the LMvehicle in the LMadapter atop the S-IVB stage. The

S-IVB stage was separated from the docked CSMand LMshortly after 4

hours into the mission.

In mannedlunar missions prior to Apollo 13, the spent S-IVB third

stages were accelerated into solar orbit by a "slingshot" maneuverin

which residual liquid oxygen was dumpedthrough the J-2 engine to provide propulsive energy. OnApollo 13, the plan was to impact the S-IVB

stage on the lunar surface in proximity to the seismometer emplaced in

the Oceanof Storms by the crew of Apollo 12.

Twohours after TLI, the S-IVB attitude thrusters were ground commandedon 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-pound stage hit the Moonlasted 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) modewhich 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 PTCto

stabilize the thermal response by spacecraft structures and systems.

At 30:40:49 g.e.t., a midcourse correction maneuverwas madeusing

the service module propulsion system. The crew preparations for the

burn and the burn itself were monitored by the Mission Control Center

(MMC)at MSCby telemetered data and by television from the spacecraft.

This midcourse correction maneuverwas 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. A similar trajectory had been

flown on Apollo 12. The objective of leaving a free-return trajectory

is to control the arrival time at the Moonto insure the proper lighting

conditions at the landing site. Apollo 8, 10, and ll flew a pure freereturn trajectory until lunar orbit insertion. The Apollo 13 hybrid

transfer maneuverlowered the predicted closest approach, or pericynthion, altitude at the Moonfrom 210 to 64 nautical miles.

From launch through the first 46 hours of the mission, the performance of oxygen tank no. 2 was normal, so far as telemetered data and

crew observations indicate. At 46:40:02, the crew turned on the fans in

oxygen tank 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. Analysis of the electrical wiring of the quantity gage

shows that this erroneous reading could be causedby either a short circuit or an open circuit in the gage wiring or a short circuit between

the gage plates. Subsequentevents indicated that a short was the more

likely failure mode.

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 (LMP)were cleared to enter the

LMto commenceinflight inspection of the LM. Groundtests before launch

had indicated the possibility of a high heat-leak rate in the LM descent

stage supercritical helium tank. Crew verification of actual pressures

found the helium pressure to be within normal limits. Supercritical

helium is stored in the LMfor pressurizing propellant tanks.

The LMwas powered downand preparations were underway to close the

LMhatch 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 CMcaution 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 MCCrequested the crew to turn

on the cryogenic system fans and heaters.

The CommandModule Pilot (CMP)acknowledgedthe 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

ILl ]LJ:- L E L: L L: L. L: u n n z L L L
About l-l/2 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. i 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). The LM, therefore, became the only source of sufficient electrical power and oxygen to permit safe return of the crew to Earth.

The various telemetered parameters of primary interest are shown

in figure 4-10 and listed in table 4-11.

4-28
Im,l Ji1_ I_ I III,I roll I" "_J I_I_ _rl:l

e',:_',', !_'_ _l::t_4 _.

N

............... ,_ _E;_ _,_

--_- t+-

+m_. i

:.i:ili:_2

::=: :4:

¥

_,_

,_td:,!

N

o o_

I I I 1 I

_. =. _. =

d

o

0

0.)

b

I1)

O)

O)

I

c;

,-4

!

.-=t

_0

oH

4-29
0_-_

AC_s I and 2v_ta_

Dc i_s A ancl II v_t_

Htiium la_ lJmi*r_url

Bay3 mWItzlr tank surfacm

t_ure

Ox_m rank I and 2 qulntlty, ten_afllt ure and pressure
TABLE 4-II.- DETAILED CHRONOLOGY FROM

2.5 MINUTES BEFORE THE ACCIDENT TO 5 MINUTES AFTER THE ACCIDENT

Time_ 6.e.t. Event

Events During 52 Seconds Prior to First Observed Abnormality

55:52:31

55:52:58

55:53:06

55:53:18

55:53:19

55:53:20

55:53:20

55:53:21

Master caution and warning triggered by low hydrogen

pressure in tank no. 1. Alarm is turned off after

4 seconds.

Ground requests tank stir.

Crew acknowledges tank stir.

Oxygen tank no. 1 fans on.

Oxygen tank no. 1 pressure decreases 8 psi.

Oxygen tank no. 2 fans turned on.

Stabilization control system electrical disturbance

indicates a power transient.

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.

55:53:22.757

55:53:22.772

1.2-volt decrease in ac bus 2 voltage.

ll.l-amp rise in fuel cell 3 current for one

sample.

55 :53:36 Oxygen tank no. 2 pressure begins rise lasting

for 24 seconds.

55:53:38.057 ll-volt decrease in ac bus 2 voltage for one

sample.

55:53:38.085 Stabilization control system electrical disturbance

indicates a power transient.

4-31
TABLE4-11.- DETAILEDCHRONOLOGY FROM

2.5 MINUTESBEFORETHEACCIDENTO 5 MINUTESAFTERTHEACCIDENT- Continued

Time, g.e.t. Event

55:53:41.172

55:53:41.192

55:54:00

55:54:15

55:54:30

55:54:31

55:54:43

55:54:45

55:54:48

55:54:51

55:54:52

55:54:52.7O3

55:54:52.763

55:54:53.182

55:5_: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.

0xygen. 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
TABLE 4-II.- DETAILED CHRONOLOGY FROM

2.5 MINUTES BEFORE 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. i pressure drops h.2 psi.

2.8-amp rise in total fuel cell current.

X, Y, and Z accelerations in CM indicate 1.17g,

0.65g and 0.65g, respectively.

55:54:53.555

55:54:53.555+

55:54:54.741

55:54:55.35

55:54:56

55:54:56

55:5h:56

55:54:56

55:54:56

1.8-Second Data Loss

Loss of telemetry begins.

Master caution and warning triggered by dc main

bus B undervoltage. Alarm is turned off in 6

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.

Recovery of telemetry data.

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

sens ors.

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.

4-33
. TABLE4-II.- DETAILEDCHRONOLOGY FROM

2.5 MINUTESBEFORETHEACCIDENTO 5 MINUTESAFTERTHEACCIDENT- Continued

Time s g.e.t. Event

55:54:56

55:54:57

55:54:59

55:55:01

55:55:02

55:55:02

55:55:09

55:55:20

55:55:35

55:55:29

55:56:10

55:56:38

Oxygen tank no. I 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 i 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
TABLE 4-II.- DETAILED CHRONOLOGY FROM

2.5 MINUTES BEFORE THE ACCIDENT TO 5 MINUTES AFTER THE ACCIDENT - Concluded

Time, g. e. t. Event

55:57:04

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

55:58:25

56:00:06

Crew reports, "That jolt must have rocked the

sensor on--see now--oxygen quantity 2. It was

oscillating down around 20 to 60 percent. Now

it's full-scale high again."

Master caution and warning triggered by dc main

bus B undervoltage. Alarm is turned off in

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.

"ac Crew reports, 2 is showing zip."

Crew reports, "Yes, we got a main bus A undervolt

now, too, showing. It's reading about 25-1/2.

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.

4-35

1/ N E '" L L L '
PART 4. SUMMARY ANALYSIS OF THE ACCIDENT

Combustion in oxygen tank no. 2 led to failure of that tank, damage

to oxygen tank no. i 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. i system integrity.

INITIATION

55:53:20*

55:53:22.757

Key Data

Oxygen tank no. 2 fans turned on.

1.2-volt decrease in ac bus 2 voltage.

*In evaluating telemetry data, consideration must be given to the

fact that the Apollo pulse code modulation (PCM) system samples data in

time and quantitizes in amplitude. For further information, reference

may be made to Part B7 of Appendix B.

4-56
55:53:22.772 ll.l-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--indicating 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 drawn from the fuel

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 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 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 arcing occurred in the fan motor or its

leads to initiate the accident sequence. The energy available from the

short circuit was probably iO 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. There

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 pointing to any other mechanism

of initiation.

PROPAGATION OF COMBUSTION

55:53:36

55:53:38.057

Key Data

Oxygen tank no. 2 pressure begins rise (same event

noted previously).

ll-volt decrease recorded in ac bus 2 voltage.

4-37

N N E 'L L L: L h L N_ 1: L,: L: Lt n J..: L .
55:53:41.172 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.

55:54:00 Oxygen tank no. 2 pressure levels off at 954 psia.

55:54:15 Oxygen tank no. 2 pressure begins to rise again.

55:54:30 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.

55:54:31 Oxygen tank no. 2 temperature begins to rise rapidly.

55:54:45 Oxygen tank no. 2 pressure reading reaches maximum

recorded value of 1008 psia.

55:54:52.763 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 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, 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 with these

4-38
observations. The cause of the 15-second period of relatively constant

pressure first indicated at 55:53:59.763 has not been precisely determined; it is believed to be associated with a changein 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.

LOSSOFOXYGENTANKNO. 2 SYSTEMINTEGRITY

55:54:52

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

Key Data

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.

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. Pressure began a decrease

for 8 seconds, dropping to 996 psia before readings were lost. Virtually

*Several bits of data have been obtained from this "loss of telemetry data" period.

4-39
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 0.22 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. When signal was reacquired, all instrument 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. Photographs taken later by the Apollo 13 crew as the SM was

jettisoned show that the bay 4 panel was ejected, undoubtedly during

this event.

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 1/2-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. If the hole formed in the pressure vessel were large

enough and formed rapidly enough, the escaping oxygen alone would be

adequate to blow off the bay 4 panel. However, it is also quite possible that the escape of oxygen 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 dome) which would augment the

4-40
4-11.-

Figure X_-_X_ Closeup view of oxygen tank shelf.

4-41

N N L. ,. ,.....
M_
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 TANK NO. i 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 to

55:54:55.35

Loss of telemetry data.

55:54:56 Oxygen tank no. 1 pressure reads 782 psia and drops

steadily. Pressure drops over a period of 130 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. Pressure decreases were recorded over a period of

approximately 130 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.

Since there is no evidence that there was any anomalous condition

arising within oxygen tank no. l, 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. l, or have caused a valve to leak because of mechanical shock.

Precedingpageblank

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 i. 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. The high-gain antenna

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 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

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 power ac bus 1 through dc

main bus A, but the failure of fuel cell 3 left dc main bus B and ac

bus 2 unpowered (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 oxygen 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.

4-4 
r

Fuel cell

1

Fuel cell

2

DC main A AC bus i l

m

it.

-LI

%I1

Fuel cell

3

J DC main B

I

Cryo 0 2

tank 2

pressure

gage

Figure

II l Inverter : 2

4-12.- Electrical configuration

Cryo 0 2

tank 2

fan

AC bus 2

I

Cryo 0 2

tank 2

quantity,

temperature

gages [

SCS

telemetry

channels

at 55:54:53 g.e.t.
The Lunar Module Pilot was in the tunnel between the CSMand the LM,

returning to the CSM. The CommandModule Pilot was in the left-hand

couch, monitoring spacecraft performance. Because of the master alarm

indicating low voltage, the CMPmovedacross to the right-hand couch

where CSMvoltages can be observed. He reported that voltages were

"looking good" at 55:56:10. At this time, main bus B had recovered and

fuel cell 3 did not fail for another l-l/2 minutes. He also reported

fluctuations in the oxygen tank no. 2 quantity, followed by a return

to the off-scale high position. (See fig. 4-13 for CMpanel arrange- ment).

Whenfuel cells 1 and 3 electrical output readings went to zero,

the ground controllers could not be certain that the cells had not somehowbeen disconnected from their respective busses and were not otherwise

all right. Attention continued to be focused on electrical problems.

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. Whenit was realized that fuel cells 1 and 3 were

not functioning, the crew was directed to perform an emergencypowerdown

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. Whenthis was done, and it

was realized that oxygen tank no. 2 had failed, the extreme seriousness

of the situation becameclear.

During the succeeding period, efforts were madeto 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 l-l/2 hours after the accident that the

oxygen tank no. 1 leak could not be stopped and that shortly it would be

necessary to use the LMas a "lifeboat" for the remainder of the mission.

By 58:40 g.e.t., the LMhad been activated, the inertial guidance

reference transferred from the CSMguidance system to the LMguidance

system, and the CSMsystems were turned off.

RETURNTOEARTH

The remainder of the mission was characteriz_ d by two main activities--planning and conducting the necessary propulsion maneuvers to

return the spacecraft to Earth, and managing the use of consumablesin

such a way that the LM, which is designed for a basic mission with two

crewmenfor a relatively short duration, could support three menand serve

as the actual control vehicle for the time required.

4-46

ILl L L
E

[

U

One significant anomaly was noted during the remainder of the

mission. At about 97 hours 14 minutes into the mission, the IAfP

reported hearing a "thump" and observing venting from the LM. 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 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 normal maneuver execution variations

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

reaction control system thrusters. All burns were satisfactory. Figures

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 for the rest

of the mission; LM oxygen for breathing; and lithium hydroxide (Li0H)

filter cannisters used to remove carbon dioxide from the spacecraft

cabin atmosphere. These consumables, 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 down to conserve electric

power and to generate less heat and thus use less water, the situation

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

crew to use materials on board to fashion a device allowing use of the

CM Li0H eannisters 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).

4-48

m ii I!

1: I

_A_J

,i*];

ILl IJ E .2L

_Ad

E

!

iJ

/

Figure 4-13.- _in display panel.

4-47 fi

r_

F_

r_

I

kO

MCC-2

for entry

corridor

(137:40)

MCC-5 ]

for entry

corridor

(105:18)

;tart of

problem

(55:55)

to

free-return

(61:30)

\

PC + 2 hr

for Pacific

landing

(79:28)

Figure 4-14.- Translunar trajectory phase.
r_

r

P_

r

r_

r

,

I

kJ3

0

Entry altitude

_ G.E.T. 142:30

E

/--- Landing

G.E.T. 142:54

CM power up

G.E.T. 140:10

LM jettison

G.E.T. 141:30

L

MCC-7

G.E.T. 137:40

LM power up

(3.E.T. 133:35

SM jettison

G.E.T. 138:01

Figure 4-15.- Final trajectory phase.
Figure 4-16.- Lithium hydroxide canister modification.

4-51
r

r

r

-kI

k.n

Po

300

250

200

100

50

0

50

Anomaly55:54

I

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II

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II

II

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entersLM (57:43)

60

I

70

Behind moon(77:09)

I

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I I I

80 90

I I

I00 Ii0

Time, hours

LM jettison(141:30)

I

I

I

I

I

I

I

I

28.2 Ibs remainingI I

120 130 140

Figure 4-17.- Usable remaining water.

150

r ..
.... j

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60

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I

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II

I

I Behind moon (77:09) I

II j

II I

II I

II j I I i _ I I

60 70 80 90 i00 II0

28.53 Ibs

remaining

LM jettison (141:30)

I

I

I

I

i i II

120 130 140

J

150

Time, hours

Figure _-19.- Usable remaining oxygen.
TABLE4-III.- CABINATMOSPHERECARBONDIOXIDE

REMOVALBYLITHIUMHYDROXIDE

Require d

Available in LM

Available in CM

85 hours

53 hours

182 hours

A more detailed recounting of the events during the Apollo 13

launch countdown and mission will be found in Appendix B to this report.

4-55
This page left blank intentionally.

NASA-- MSC-- Coral., Houston, Texas

4-56
CHAPTER 5

FINDINGS, DETERMINATIONS, AND RECOMMENDATIONS

5-0

PART I. INTRODUCTION

The following findings, determinations, and recommendations are the

product of about 7 weeks of concentrated review of the Apollo 13 accident

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

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.

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 electrically initiated fire in oxygen

tank no. 2 in the service module (SM) was the cause of the accident. Accordingly, the Board 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 accident is generally understood, and

the most probable cause has been identified. However, at the time of this

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

Further tests and analyses, which will be carried out under the overall direction of MSC, will continue to generate new information relative

to this accident. It is possible that this evidence may lead to conclusions differing in detail from those which can be drawn now. However, it

is most unlikely that fundamentally 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 organizational elements responsible for it. It was found that the accident was

not the result of a chance malfunction in a statistical sense, but rather

resulted from an 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, which were inadequate and

would subsequently fail during ground test operations at Kennedy Space

Center (KSC).

5-1
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 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 the displaced fill tube assembly.

f. The special detanking procedures at KSCsubjected the tank to an

extended period of heater operation and pressure cycling. Theseprocedures 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

NASA,NR, and Beech did not recognize the possibility of damagedue to

overheating. Manyof these officials were not aware of the extended

heater operation. In any event, adequate thermostatic switches might

have been expected to protect the tank.

h. A numberof factors contributed to the presence of inadequate

thermostatic switches in the heater assembly. The original 1962 specifications from NRto BeechAircraft Corporation for the tank and heater

assembly specified the use of 28 V dc power, which is used in the spacecraft. In 1965, NRissued a revised specification which stated that the

heaters should use a 65 V dc power supply for tank pressurization; this

was the power supply used at KSCto reduce pressurization time. Beech

ordered switches for the Block II tanks but did not change the switch

specifications to be compatible with 65 V dc.

i. The thermostatic switch discrepancy wasnot detected by NASA,NR,

or Beech in their review of documentation, nor did tests identify the incompatibility of the switches with the ground support equipment (GSE)at

KSC, since neither qualification nor acceptance testing required switch

cycling under load as should have been done. It was a serious oversight

in which all parties shared.

j. The thermostatic switches could accommodatethe 65 V dc during

tank pressurization because they normally remained cool and closed. However, they could not open without damagewith 65 V dc power applied. They

were never required to do so until the special detanking. During this

5-2
procedure, as the switches started to open whenthey 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 KSCif 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 have beenexpected.

i. As shownby subsequent tests, failure of the thermostatic switches

probably permitted the temperature of the heater tube assembly to reach

about i000 ° F in spots during the continuous 8-hour period of heater

operation. Suchheating has been shownby tests to severely damagethe

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 whenfilled with oxygen and electrically powered.

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

the fan motor wiring, possibly movedby the fan stirring, short circuited

and ignited its insulation by meansof 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 augmentedby 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. i, damagedthe 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

membersof 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 managementlessons

which can be applied to help assure the success of future space flight

missions; several recommendationsof this nature are included.

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

a critical nature. The report highlights in detail faults or deficiencies

in equipment and procedures that the Board has identified. This is the

nature of a review board report.

5-3
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. It has flown to the Moon five times and landed

twice. The tank which failed_ the design of which is criticized in this

report, is one of a series which had thousands of hours of successful

operation in space prior to Apollo 13.

While the team of designers, engineers_ and technicians that build

and operate the Apollo spacecraft also has shortcomings, the accomplishments speak for themselves. By hardheaded self-criticism and continued

dedication_ this team can maintain this nation's preeminence in space.

5-4
PART 2. ASSESSMENT OF ACCIDENT

FAILURE OF OXYGEN TANK NO. 2

l. Findings

a. 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. i, and

a resulting loss of power from the oxygen-fed fuel cells.

b. There is no evidence of any forces external to oxygen tank

no. 2 during the flight which might have caused its failure.

Co Oxygen tank no. 2 contained materials, including Teflon and

aluminum, which if ignited will burn in supercritical

oxygen.

d. Oxygen tank no. 2 contained potential ignition sources:

electrical wiring, unsealed electric motors, and rotating

aluminum fans.

e. During the special detanking of oxygen tank no. 2 following

the countdown demonstration test (CDDT) at KSC, the thermostatic switches on the heaters were required to open while

powered by 65 V dc in order to protect the heaters from overheating. The switches were only rated at 30 V dc and have

been shown to weld closed at the higher voltage.

f. Data indicate that in flight the tank heaters located in

oxygen tanks no. i and no. 2 operated normally prior to the

accident, and they were not on at the time of the accident.

g° 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.

h. Telemetry data izmnediately 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.

i. The pressure and temperature within oxygen tank no. 2 rose

abnormally during the 1-1/2 minutes immediately prior to the

accident.

5-5
Determinations

(i) The cause of the failure of oxygen tank no. 2 was combustion

within the tank.

(2) Analysis showed that the electrical energy flowing into the

tank could not account for the observed increases in pressure

and temperature.

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

initiate the accident sequence.

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

of Teflon wire insulation on the fan motor wires, caused by

electric arcs in this wiring.

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.

(6) The telemetered data indicated electrical arcs of sufficient

energy to ignite the Teflon insulation, as verified by subsequent tests. These 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.

(7) 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 demonstrated by subsequent tests.

(8) The 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.

SECONDARY EFFECTS OF TANK FAILURE

o Findings

a. Failure of the tank was accompanied by several events including:

5-6
A %ang" 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).

Momentary loss of telemetry.

Closing of several valves by shock loading.

Loss of integrity of the oxygen tank no. i 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.

b ° The panel covering of bay 4 could be blown off by pressurization of the bay. About 25 psi of uniform pressure in bay 4

is required to blow off the panel.

Co The various bays and sectors of the SM are interconnected

with 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 i0 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. This possibility

was substantiated by a series of special tests.

(2) 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 oxygen and hot particles

emerging from the top of the tank, as demonstrated in subsequenttests.

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.

(4) Photographs of the SM by the crew did not establish the

condition of the oxygen tank no. 2.

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

by the panel_ or a portion thereof, as it left the SM.

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

loss of power resulted from the tank no. 2 failure.

(7)

(s)

Telemetry, although good, is insufficient to pin down the

exact nature, sequence, and location of each event of the

accident in detail.

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 TANK 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

system.

c. The heaters, fans, and tank instrumentation are used in the

measurement and management of the oxygen supply.

Determinations

(i) The storage of supercritical oxygen was appropriate for the

Apollo system.

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

supply is used.

(3) Fans were used to prevent excessive pressure drops due to

stratification, to mix the oxygen to improve accuracy of

5-8
(4)

(5)

quantity measurements, and to insure adequate heater input

at low densities and high oxygen utilization rates. The

need for oxygen stirring on future flights requires further

investigation.

The amount of material in the tank which could be ignited

and burned in the given environment could have been reduced

significantly.

The potential ignition sources constituted an undue hazard

when considered in the light of the particular tank design

with its assembly difficulties.

(6) NASA, the prime contractor, and the supplier of the tank

were not fully aware of the extent of this hazard.

(7) Examination of the high-pressure oxygen system in the service

module following the Apollo 204 fire, which directed attention to the danger of fire in a pure oxygen environment,

failed to recognize the deficiencies of the tank.

PREFLIGHT DAMAGE TO TANK WIRING

0 Findings

a. The oxygen tank no. 2 heater assembly contained two thermostatic switches designed to protect the heaters from overheating.

b. The thermostatic switches were designed to open and interrupt

the heater current at 80 ° ± i0 ° F.

c. The heaters are operated on 28 V dc in flight and at NR.

d. The heaters are operated on 65 V ac at Beech Aircraft Corporation and 65 V dc at the Kennedy Space Center. These

higher voltages are used to accelerate tank pressurization.

e. The thermostatic switches were rated at 7 amps at 30 V dc.

While they would carry this current at 65 V dc in a closed

position, they would fail if they started to open to interrupt this load.

fo Neither qualification nor acceptance testing of the heater

assemblies or the tanks required thermostatic switch opening

to be checked at 65 V dc. The only test of switch opening

5-9
was a continuity check at Beech in which the switch was

cycled open and closed in an oven.

g. The thermostatic switches had never operated in flight because this would only happen if the oxygen supply in a tank

were depleted to nearly zero.

hi 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 switches cool and closed.

i. During the CDDT, the oxygen tank no. 2 would not detank in

a normal manner. On March 27 and 28, a special detanking

procedure was followed which subjected the heater to about

8 hours of continuous operation until the tanks were nearly

depleted of oxygen.

j. A second special detanking of shorter duration followed on

March 30, 1970.

k. The oxygen tanks had not been qualification tested for the

conditions encountered in this procedure. However, specified allowable heater voltages and currents were not exceeded.

i. 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.

m. The oxygen tank heater controls at KSC contained ammeters

which would have indicated thermostatic switch operation.

Determinations

(i) 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

powered by 65 V dc and were probably welded shut.

(2) 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 the heaters had

reached their temperature limit and switch opening should

have been expected.

5-10
(3)

(4)

(5)

(6)

The fact that the switches were not rated to open at 65 V dc

was not detected by NASA, NR, or Beech in their reviews of

documentation or in qualification and acceptance testing.

The failed switches resulted in severe overheating. Subsequent tests showed that heater assembly temperatures could

have reached about i000 ° F.

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.

5-ii
PART 3. SUPPORTING CONSIDERATIONS

DESIGN, MANUFACTURING, AND TEST

5. Finding

The pressure vessel of the supercritical oxygen tank is constructed 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.

Co 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

(i) The stator winding insulation is brittle and easily fractured

during manufacture of the stator coils.

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

7. Findings

a. The cryogenic oxygen storage tanks contained materials that

could be ignited and which will burn under the conditions

5-12
prevailing within the tank, including Teflon, aluminum,

solder, and Drilube 822.

b. The tank contained electrical wiring exposed to the supercritical oxygen. The wiring was insulated with Teflon.

c. Somewiring was in close proximity to heater elements and

to the rotating fan.

do The design was such that the assembly of the equipment was

essentially "blind" and not amenable to inspection after

completion.

e. 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.

f. Portions of this wiring remained unsupported in the tank on

completion of assembly.

Determinations

(i) The tank contained a hazardous combination of materials and

potential ignition sources.

(2) Scraping of the electrical wiring insulation against metal

inner parts of the tank constituted a substantial cumulative

hazard during assembly, handling, test, checkout, and operational use.

(5) "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.

(4) 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.

(5) The design was such that it was difficult to insure against

these hazards.

(6) There is no evidence that the wiring was damaged during manufacturing.

5-13
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.

c. Probing with a hand tool was used in manufacturing to compensate for limited visibility of the tube segment positions.

Determination

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

relatively loose fill tube parts that could go undetected in

final inspection and be subsequently displaced.

i0. Findings

a. The 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. One 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 lO-amp circuit breaker.

Determination

The fuel cell oxygen supply valve module has been identified as

potentially hazardous.

ll. 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 SM.

5-14
b. The SMundergoes 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.

C. The oxygen shelf assembly incorporating S/N XTAO008 in the

tank no. 2 position, which had been shaken during removal

from SM 106, was installed in SM 109 one month later.

d. An analysis, shelf inspection, and a partial retest emphasizing electrical continuity of internal wiring were accomplished before reinstallation.

Determinations

(i) Displacement of fill tube parts could have occurred, during

the "shelf drop" incident at the prime contractor's plant,

without detection.

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

but special tests and analyses indicate that this is unlikely.

(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
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.

b, A special detanking procedure was used to empty oxygen tank

no. 2 after CDDT. This procedure involved continuous protracted heating with repeated cycles of pressurization to

about 300 psi with warm gas followed by venting.

c. It was employed both after CDDT and after a special test to

verify that the tank could be filled.

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.

fo 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 i000 ° F range would

exist at locations on the heater tube assembly that were in

close proximity with the motor wires. These temperatures

are high enough to damage Teflon and melt solder.

Determinations

(i) 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.

(2) Such evidence indicates that the tank had undergone some

change of internal configuration during the intervening

events of the previous 3 years.

5-16
(3) 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 i000 ° F.

(4) Severe damage to the insulation of electrical wiring internal

to the tank, as determined from subsequent tests, resulted

from the special procedure.

(5) Damage to the insulation, particularly on the long unsupported lengths of wiring, may also have occurred due to

boiling associated with this procedure.

(6) MSC, KSC, and NR 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. They made appropriate contacts before

making the 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

(l) NR and MSC personnel who prepared the TSCD did not know that

the tank heater thermostatic switches would not protect

the tank.

(2) Launch operations personnel assumed the tank was protected

from overheating by the switches.

5-17
16.

b) Launch operations personnel at KSC stayed within the

specified tank heater voltage and current limits during the

detanking at KSC.

Findings

ao After receipt of the Block II oxygen tank specifications

from NR, 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

(i) NR and Beech specifications governing the powering and the

thermostatic switch protection of the heater assemblies were

inadequate.

(2) The specifications governing the testing of the heater

assemblies were inadequate.

17. Finding

The hazard associated with the long heater cycle during detanking

was not given consideration in the decision to fly oxygen tank

no. 2.

Determinations

(i)

(2)

MSC, KSC, and NR personnel did not know that the tank heater

thermostatic switches did not protect the tank from overheating.

If the long period of continuous heater operation with failed

thermostatic switches had been known, the tank would have

been replaced.

5-18

L . z,.-,,_ A-_
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.

b° When the Apollo 13 cryogenic oxygen system was originally

designed, the management controls were not defined in as

great detail as they are now.

Determination

From review of documents and interviews, it appears that the

management controls existing at that time were adhered to in

the case of the cryogenic oxygen system incorporated in

Apollo 13.

19. Finding

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 THI_OUGH ACCIDENT

20. Findings

a. The center engine of the S-If 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 O.ig vibration in the CM.

Determinations

(i) Investigation of this S-If anomaly was not within the purview

of the Board except insofar as it relates to the Apollo 13

accident.

5-19
(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. i 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

(i) The oxygen tank no. 2 quantity probe short circuited at

46:40:08.

(2) The short circuit could have been caused by either a completely loose fill tube part or a solder splash being carried

by the moving fluid into contact with both elements of the

probe capacitor.

22. Findings

a. The crew acknowledged Mission Control's request to turn on

the tank fans at 55:53:06.

b. Spacecraft current increased by i ampere at 55:53:19.

c. The oxygen tank no. i pressure decreased 8 psi at 55:53:19

due to normal destratification.

Determination

The fans in oxygen tank no. i 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.

b° Stabilization and Control System (SCS) gimbal command telemetry channels, which are sensitive indicators of electrical

transients associated with switching on or off of certain

spacecraft electrical loads, showed a negative initial transient during oxygen tank no. 2 fan turnon cycles and a positive initial transient during oxygen tank no. 2 fan turnoff

5-20
cycles during the Apollo 13 mission. A negative initial

transient was measuredin the SCSat 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

(i)

(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 ll.l-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

(i) A short circuit occurred in the circuits of the fans in

oxygen tank no. 2 which resulted in either blown fuses or

opened wiring, and one fan ceased to function.

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

of i0 joules which, as shown in subsequent tests, is more

than sufficient to ignite Teflon wire insulation by means

of an electric arc.

25. Findings

a. A momentary ll-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.

C. After the electrical transients, CM current and ac bus 2

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

Determination

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

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

954 psia between 55:53:36 and 55:54:00. It then remained nearly

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.

Determinations

(i)

(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 i000 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

28.

This drop resulted from the normal operation of the pressure

relief valve as verified in subsequent tests.

Findings

ao 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.

Co The low-pressure trip level of the master caution and warning

system for the cryogenic storage system is only i psi below

the specified lower limit of the pressure switch which controls the tank heaters. A small imbalance in hydrogen tank

5-22
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 it

did come on, Mission Control did not observe it.

Determinations

(i) If the pressure switch setting and master caution and warning

trip levels were separated by a greater pressure differential,

there would be less likelihood of unnecessary master alarms.

(2) With the present master caution and warning system, a spacecraft problem can go unnoticed because of the presence of a

previous out-of-tolerance condition in the same subsystem.

Although a master alarm at 55:54:29 or observance of a limit

sense light in Mission Control could have alerted the crew

or Mission Control in sufficient time to detect the pressure

rise in oxygen tank no. 2_ no action could have been taken

at that time to prevent the tank failure. However_ the information could have been helpful to Mission Control and the

crew in diagnosis of spacecraft malfunctions.

(4) The limit sense system in Mission Control can be modified to

constitute a more positive backup warning system.

29. Finding

Oxygen tank no. 2 telemetry showed a temperature rise of 38 ° F

beginning at 55:54:31 sensed by a single sensor which measured

local temperature. This sensor indicated off-scale low at

55:54:53.

Determinations

(i) An abnormal and sudden temperature rise occurred in oxygen

tank no. 2 at approximately 55:54:31.

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

had progressed to the vicinity of the sensor.

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

5-23
30. Finding

Oxygen tank no. 2 telemetry indicated the following changes:

(i) 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

(i) 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 i00 times per second, began

indicating spacecraft motions. These disturbances were

erratic, but reached peak values of 1.17g, 0.65g, and 0.65g

in the X, Y, and Z 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.

d° Doppler tracking data measured an incremental velocity component of 0.26 fps along a line from the Earth to the spacecraft at approximately 55:54:55.

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.

g. Crew observations and photographs showed the bay 4 panel to

be missing and the high-gain antenna to be damaged.

5-24

ILl II E E E E k L '" " ' "
Determinations

(i) The spacecraft was subjected to abnormal forces at approximately 55:54:53. These disturbances were reactions resulting

from failure and venting of the oxygen tank no. 2 system and

subsequent separation and ejection of the bay 4 panel.

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

section thereof from bay 4 at the time of panel separation.

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 plac e in the SM at approximately the time of panel

separation.

33. Findings

a. The telemetered nitrogen pressure in fuel cell i was offscale low at reacquisition of data at 55:54:55.

b. Fuel cell I continued to operate for about 3 minutes past

this time.

Co The wiring to the nitrogen sensor passes along the top of

the shelf which supports the fuel cells immediately above

the oxygen tanks.

Determinations

(i) The nitrogen pressure sensor in fuel cell i or its wiring

failed at the time of the accident.

(2) The failure was probably caused by physical damage to the

sensor wiring or shock.

(3) This is the only known instrumentation failure outside the

oxygen system at that time.

34. Finding

Oxygen tank no. i 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
Determination

A leak caused loss of oxygen from tank no. i beginning at approximately 55:54:54.

35. Findings

a. Oxygen flow rates to fuel cells I 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.

b° 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 magnitudeas

the shock required to close the RCS valves.

c. Fuel cells i and 3 failed at about 55:58.

Determination

The oxygen supply valves to fuel cells i 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. This was quickly

followed by the crew's report of a problem.

b. 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:

5-26
Oxygentank no. i: Pressure: Several hundred psi below

normal

Quantity: Normal

Temperature: Normal

Oxygentank no. 2: Pressure: Off-scale low

Quantity: Off-scale high

Temperature: Off-scale high

C. The nitrogen pressure in fuel cell i indicated zero, which was

incompatible with the hydrogen 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 indications could have been a cryogenic storage system

instrumentation failure, and began pursuing this line of investigation.

Determination

Under these 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 i and 3 were less than 0.i ib/hr. These were unusually

low readings for the current being drawn.

b. Fuel cells i and 3 failed at about 3 minutes after the data

loss.

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.

d. Near-zero oxygen flow to fuel cells i and 3 was noted after

the main bus B failure, but this was consistent with no power

output from the fuel cells.

e . The flight controllers believed that the fuel cells could

have been disconnected from the busses and directed the crew

to connect fuel cell i 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.

(2) 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.

38. Finding

The fuel cell reactant'valve talkback indicators in the spacecraft do not indicate closed unless both the hydrogen and oxygen

valves are closed.

Determinations

(l) If these talkbacks 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 fuel cells i and 3, although they would nevertheless have failed when oxygen tank no. i ceased to supply

oxygen.

5-28
(2) The ultimate outcome would not have been changed, but had the

fuel cells not failed, Mission Control and the crew would not

have mad 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. 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. The secondary fuel

pressurization valves for quads A and C also were reported

closed.

b° 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.

C° During the l-i/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.

d. Both the valve solenoids and the onboard indications of valve

position of the propellant isolation valves for quad C are

powered by dc main bus B.

e ° During the l-i/2-hour period immediately following the

accident, Mission Control advised the crew which SM RCS

thrusters to power and which ones to unpower.

Determinations

(i) 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.

(3) Mission Control correctly determined the status of the RCS

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

5-29
Managementof Electrical System

40. Findings

a. After fuel cell I 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 o 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

(i) 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.

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

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
b. When these procedures failed to arrest the oxygen loss,

Mission Control directed the crew to shut down fuel cells i

and 3 by closing the hydrogen and oxygen flow valves.

Determinations

(i) Under more normal conditions oxygen pressure might have been

increased by turning on heaters and fans in the oxygen tanks;

no other known actions had such a possibility.

(2) There was a possibility that oxygen was leaking downstream

of the valves; had this been true, closing of the valves

might have preserved the remaining oxygen in oxygen tank

no. i.

Lunar Module Activation

42. Findings

a° With imminent loss of oxygen from oxygen tanks no. i 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. Mission Control and the crew delayed LM activation until

about 15 minutes before the SM oxygen supply was depleted.

C° There were three different LM activation checklists contained

in the Flight Data File for normal and contingency situations;

however, none of these was appropriate for the existing situation. It was necessary to activate the LM as rapidly as

possible to conserve LM consumables and CM reentry batteries

to the maximum extent possible.

d. Mission Control modified the normal LM activation checklist

and referred the crew to specific pages and instructions.

This bypassed unnecessary steps and reduced the activation

time to less than an hour.

e ° The LM inertial platform was aligned during an onboard checklist procedure which manually transferred the CM alignment to

the LM.

5-31
Determinations

45.

(i) Initiation of LM activation was not undertaken sooner because

the crew was properly more concerned with attempts to conserve

remaining SM oxygen.

(2) Mission Control was able to make workable on-the-spot modifications to the checklists which sufficiently shortened the

time normally required for powering up the LM.

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.

bo This caused some concern in Mission Control because of the

possibility of the spacecraft drifting into inertial platform

gimbal lock condition.

C. The Command Module Pilot (CMP) stated that he was not concerned because he could have quickly reestablished direct

manual attitude control if it became necessary.

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. LM flight controllers were on duty in Mission Control at the

time of the accident in support of the scheduled crew entry

into the LM.

b° If the accident had occurred at some other time during the

translunar coast phase, LM system specialists would not have

been on duty, and it would have taken at least 30 minutes to

get a fully manned team in Mission Control.

Determination

Although LM flight controllers were not required until more than

an hour after the accident, it was beneficial for them to be

present as the problem developed.

5-52
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.

C° 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.

d. Mission Control developed contingency plans for further reduction of LM power for use in case an LMbattery 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.

f. _hen 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

Determinations

(i) Earlier contingency plans and available checklists were

adequate to extend life support capability of the LM well

beyond its normal intended capability.

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

further increase the LM consumables margins.

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 i00 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.

do The crew made the modification at 93 hours, and carbon

dioxide partial pressure in the LM dropped rapidly from

7.5mm Hg to O.imm 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, LM

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
b. At about that time the Lunar Module Pilot (LMP) heard a

"thump" from the vicinity of the LM descent stage.

c. When 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 i and 2 are located. This venting continued for a few minutes.

d. Prior to 97:13 the battery load-sharing among the four

batteries had been equal, but immediately after the battery

currents returned to nominal, batteries i and 2 supplied 9

of the ii 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.

Determinations

(I) An anomalous incident occurred in the LM electrical system

at about 97:13:53 which appeared to be a short circuit.

(2) The thump and the venting were related to this anomaly.

(9) The apparent short circuit cleared itself.

(4) This anomaly was not directly related to the CSM or to the

accident.

(5) This anomaly represents a potentially serious electrical

problem.

CM Battery Recharging

48. Findings

a. About one half of the electrical capacity of reentry

battery A (20 of 40 amp-hours) was used during emergency

conditions following the accident. A small part of the

capacity of reentry battery B was used in checking out dc

main bus B at 95 hours. The reduced charge remaining in the

batteries limited the amount of time the CM could operate

after separation from the LM.

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 154 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 Changes For Safe Return to Earth

49. Findings

a° After the accident, it became apparent that the lunar landing

could not be accomplished and that the spacecraft trajectory

must be altered for a return to Earth.

b. At the time of th@ accident, the spacecraft trajectory was

one which would have returned it to the vicinity of the Earth,

but it would have been left in orbit about the Earth rather

than reentering 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 (DPS), required to return the spacecraft to the Earth.

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

the DPS engine, to shift the landing point from the Indian

Ocean to the Pacific and to shorten the return trip by

9 hours.

A 7.8-fps burn at 105:18 using the DPS engine to lower Earth

perigee from 87 miles to 21 miles.

5-36
A 3.2-fps correction at 137:40 using LM RCSthrusters, to

assure that the CMwould reenter the Earth's atmosphere at

the center of its corridor.

d. 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.

e. Without the CM guidance and navigation system, the crew could

not navigate or compute return-to-Earth maneuver target parameters.

Determinations

(1) This series of course corrections was logical and had the

best chance of success because_ as compared to other options_

it avoided use of the damaged 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.

(2) Mission Control trajectory planning and maneuver targeting

were essential for the safe return of the crew.

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.

5-37
et The reentry preparation time had to be extended in order to

accomplish the additional steps required by the unusual situation.

fe

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.

g. The procedures for accomplishing the final course correction

and the reentry preparation were developed by operations

support personnel under the direction of Mission Control.

h°

An initial set of procedures was defined within 12 hours

after the accident• These were refined and modified during

the following 2 days, and evaluated in simulators at MSC and

KSC by members of the backup crew.

i. 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

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

(i) 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 long training period.

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

51. Findings

ao

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.

5-38
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 CM 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

i. 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

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. The warning systems on board the Apollo spacecraft and in the Mission

Control Center should be carefully reviewed and modified where appropriate,

with specific attention to the following:

a. Increasing the differential between master alarm trip levels and

expected normal operating ranges to avoid unnecessary alarms.

b. Changing the caution and warning system logic to prevent an outof-limits alarm from blocking another alarm when a second quantity in the

same subsystem goes out of limits.

c. Establishing a second level of limit sensing in Mission Control

on critical quantities with a visual or audible alarm which cannot be

easily overlooked.

d. Providing independent talkback indicators for each of the six

fuel cell reactant valves plus a master alarm when any valve closes.

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

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

analyses now underway in order to understand more compl@tely the details

of the Apollo 13 accident. In addition, the lunar module power system

anomalies should receive careful attention. Other NASA Centers should

continue their support to MSC in the areas of analysis and test.

5-40
6. Wheneversignificant anomalies occur in critical subsystems during

final preparation for launch, standard procedures should require a presentation of all prior anomalies on that particular piece of equipment, including those which have previously been corrected or explained. Furthermore, critical decisions involving the flightworthiness of subsystems

should require the presence and full participation of an expert who is

intimately familiar with the details of that subsystem.

7. NASAshould 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. NASAshould conduct additional research on materials compatibility,

ignition, and combustion in strong oxidizers at various g levels; and on

the characteristics of supercritical fluids. Whereappropriate, new NASA

design standards should be developed.

9. The MannedSpacecraft Center should reassess all Apollo spacecraft

subsystems, and the engineering organizations responsible for them at

MSCand 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. Wherenecessary, 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.

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NASA -- MSC