REPORT OF APOLLO 13 REVIEW BOARD

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

APOLLO 13 REVIEW BOARD

The Honorable Thomas 0. Paine
Administrator
National Aeronautics and Space Administration
Washington, D.C. 20546
June 15, 1970

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

Preface

     The Apollo 13 accident, which aborted man's third mission to explore
the surface of the Moon, is a harsh reminder of the immense difficulty of
this undertaking.

     The total Apollo system of ground complexes, launch vehicle, and
spacecraft constitutes the most ambitious and demanding engineering
development ever undertaken by man. For these missions to succeed, both
men and equipment must perform to near perfection.  That this system has
already resulted in two successful lunar surface explorations is a tribute
to those men and women who conceived, designed, built, and flew it.

     Perfection is not only difficult to achieve, but difficult to
maintain.  The imperfection in Apollo 13 constituted a near disaster,
averted only by outstanding performance on the part of the 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 made effective 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 manned space flight.

TABLE OF CONTENTS

LETTER OF TRANSMITTAL
PREFACE
TABLE OF CONTENTS

CHAPTER 1 - 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
Manned Space Flight

NASA Management Instruction 8621.1, Subject:
Mission Failure Investigation Policy and Procedures,
April 14, 1966

NASA Management Instruction 1156.14, Subject:
Aerospace Safety Advisory Panel, December 7, 1967 . . .

CHAPTER 2 - BOARD HISTORY AND PROCEDURES

Part 1. Summary of 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

CHAPTER 3 - DESCRIPTION OF APOLLO 13 SPACE VEHICLE AND
MISSION

Part 1. Apollo/Saturn V Space Vehicle
Part 2. Apollo 13 Mission Description

Chapter 4 - REVIEW AND ANALYSIS OF APOLLO 13 ACCIDENT
Part 1. Introduction
Part 2. Oxygen Tank No. History
Part 3. Apollo 13 Flight
Part 4. Summary Analysis of the Accident
Part 5. Apollo 13 Recovery

CHAPTER 5 - FINDINGS, DETERMINATIONS, AND RECOMMENDATIONS
Part 1. Introduction
Part 2. Assessment of Accident
Part 3. Supporting Considerations
Part 4. Recommendations

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

APPENDIX B - REPORT OF MISSION EVENTS PANEL

APPENDIX C - REPORT OF MANUFACTURING AND TEST PANEL

APPENDIX D - REPORT OF DESIGN PANEL

APPENDIX E - REPORT OF PROJECT MANAGEMENT PANEL

APPENDIX F - SPECIAL TESTS AND ANALYSES

APPENDIX G - BOARD ADMINISTRATIVE PROCEDURES

APPENDIX H - BOARD RELEASES AND PRESS STATEMENTS

CHAPTER 1
AUTHORITIES

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

1. It is NASA policy as stated in Reference (a) "to investigate and
document the causes of all major mission failures which occur in the
conduct of 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

     (c) Direct such further specific investigations as may be necessary.

     (d) Report as soon as possible its findings relating to the cause or
causes of the accident and the effectiveness of the flight and ground
recovery actions.

     (e) Develop recommendations for corrective or other actions, based
upon its findings and determinations or conclusions derived
therefrom.

     (f) Document its findings, determinations, and recommendations and
submit a final report.

4. As Chairman of the Board you are delegated the following powers:

     (a) To establish such procedures for the organization and operation
of the Board as you find most effective; such procedures shall be part of
the Board's records. The procedures shall be furnished the Aerospace
Safety Advisory Panel for its review and comment.

     (b) To establish procedures to assure the execution of your
responsibilities in your absence.

     (c) To designate such representatives, consultants, experts, liaison
officers, observers, or other individuals as required to support the
activities of the Board. You shall define their duties and
responsibilities as part of the Board's records.

     (d) To keep us advised periodically concerning the organization,
procedures, operations of the Board and its associated activities.

5. By separate action we are requesting the Aerospace Safety Advisory
Panel established by Reference (b) to review both the procedures and
findings of the Board and submit its independent report to us.

6. By separate action we are directing the Associate Administrator for
Manned Space Flight to:

     (a) Assure that all elements of the Office of Manned Space Flight
cooperate fully with the Board and provide records, data, and technical
support as requested.

     (b) Undertake through the regular OMSF organization such reviews,
studies, and supporting actions as are required to develop 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.

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

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

WASHINGTON. D.C.     20546

Office of the Administrator                          April 21, 1970

TO: Mr. Edgar M. Cortright

SUBJECT: Membership of Apollo 13 Review Board

Reference: Memorandum to you of April 17, subject: Establishment of
Apollo 13 Review Board

In accordance with paragraph 2(a) of Reference (a), the membership of the
Apollo 13 Review Board is established as follows:

Members:

Mr. Edgar M. Cortright, Chairman (Director, Langley Research Center)
Mr. Robert F. Allnutt (Assistant to the Administrator, NASA Hqs.)
Mr. Neil Armstrong (Astronaut, Manned Spacecraft Center)
Dr. John F. Clark (Director, Goddard Space Flight Center)
Brig. General Walter R. Hedrick, Jr. (Director of Space, DCS/RED,
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)

OMSF Technical Support:

Mr. Charles W. Mathews (Deputy Associate Administrator, Office of
Manned Space Flight)

Observers:

Mr. William A. Anders (Executive Secretary, National Aeronautics
and Space Council)

Dr. Charles D. Harrington (Chairman, NASA Aerospace Safety
Advisory Panel)
Mr. I. I. Pinkel (Director, Aerospace Safety Research and
Data Institute, Lewis Research Center)

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

     NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
     WASHINGTON, D.C. 20546

April 20, 1970

Office of the Administrator

T0 : Dr. Charles D. Harrington
Chairman, Aerospace Safety Advisory Panel

SUBJECT:  Review of Procedures and Findings of Apollo 13 Review
Board

Attachment: (a) Memorandum dated April 17, 1970, to Mr. Edgar M.
Cortright,
Subject: Establishment of Apollo 13 Review Board

References: (a) Section 6, National Aeronautics and Space Administration
                Authorization Act, 1968

            (b) NMI 1156.14 - Aerospace Safety Advisory Panel

1. In accordance with References (a) and (b), the Aerospace Safety
Advisory Panel (hereafter referred to as the Panel) is requested to review
the procedures and findings of the Apollo 13 Review Board (hereafter
referred to as the Board) established by Attachment (a).

2. The procedures established by the Board will be made available to the
Panel for review and comment as provided in paragraph 4(a) of Attachment
(a).

3. As Chairman of 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 10 days after the submission to us of the final report
of the Board.

George M. Low, Deputy Administrator    T. O. Paine, Administrator

Enclosure
cc: Mr. Edgar M. Cortright, Chairman, Apollo 13 Review Board
M/Mr. Dale Myers

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
WASHINGTON, D.C 20546

April 20, 1970
OFFICE or THE ADMINISTRATOR

TO : Mr. Dale D. Myers

Associate Administrator for Manned Space Flight

SUBJECT : Apollo 13 Review

References: (a) Memorandum dated April 17, 1970, to Mr. Edgar M.
Cortright,
subject: Establishment of Apollo 13 Review Board

(b) Memorandum dated April 20, 1970, to Dr. Charles D.
Harrington, subject: Review of Procedures and Findings
of Apollo 13 Review Board

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

(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

NMI  8621.1

April 14, 1966

Management Instruction
SUBJECT: MISSION FAILURE INVESTIGATION POLICY AND
PROCEDURES

1. PURPOSE

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

This Instruction is applicable to NASA Headquarters and field
installations.

3. DEFINITION

For the purpose of this Instruction, the following term shall apply:

Instruction: Failure is defined as not achieving a major mission
objective.

4. POLICY

a. It is NASA policy to investigate and document the causes of all major
mission failures which occur in the conduct of its space and aeronautical
activities and to take appropriate corrective actions as a result of the
findings and recommendations.

b. The Deputy Administrator may conduct independent investigations of
major failures in addition to those investigations required of the
Officials-in-Charge of Headquarters Program Offices as set forth in
paragraph 5a

5. PROCEDURES

a. Officials-in-Charge of Headquarters Program Offices are responsible,
within their assigned areas, for:

(1) Informing promptly the Deputy Administrator of each major
failure and apprising him of the nature of the failure, status
of investigations, and corrective or other actions which are
or will be taken.

(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.

b. When the Deputy Administrator decides to conduct an independent
investigation, he will:

(l) 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 Boards
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) Documenting findings, determinations and recommendations for
corrective or other actions and submitting such
documentation to the Deputy Administrator.

c. Procedures for implementing the Board's recommendations shall be
determined by the Deputy Administrator.

6 CANCELLATION

NASA Management Manual Instruction 4-1-7 (T.S. 760), March 24,
l964.

Deputy Administrator

DISTRIBUTION:
SDL l

NMI 1156.14

December 7, 1967
Effective Date

Management  Instruction
SUBJECT: AEROSPACE SAFETY ADVISORY PANEL

1. 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. I68, 170). Since the Panel was established by
statute, its formation and use are not subject to the provisions of
Executive Order 11007 or of NMI 1150.2, except to the extent that
such provisions are made applicable to the Panel under this
Instruction.

3. DUTIES

a. The duties of the Panel are set forth in Section 6 of the
National Aeronautics and Space Administration Authorization
Act, 1968, as follows:

"The Panel shall review safety studies and operations plans referred to it
and shall 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, and public safety activities, and the management of
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.

(d) Establishment and implementation of safety standards
and procedures for operation of program support and
administrative aircraft.

(2) Systems Safety. This element includes those activities
specifically organized to deal with the potential hazards of
complex 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.

(d) Establishment and implementation of systems safety
plans.

(3) 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
NASA's business, but which may nevertheless be exposed to
potential hazards associated with carrying 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.

(4) 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 follow-up on
corrective actions and the feedback of accident/incident
information to other involved or interested
organizations.

(d) The analysis and dissemination of safety data.

4. 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 discharge
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 planned
and on-going activities in this system.

5. ORGANIZATION

a. Membership

(1) 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 Authorization Act,
1968.

b. Officers

(1) 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.

c. Committees

(1) 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 of the Panel.

(2) All such committee 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) 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.

(4) 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.

(5) 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.

d. Meetings

(1) Regular meetings of the Panel will be held as often as
necessary and at least twice a year. One meeting each year
shall be an Annual Meeting. Business conducted at this
meeting will include selecting the Chairman and the Vice
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.

(2) Special meetings of the Panel may be called by the Chairman,
by notice served personally upon or by mail or telegraph to
the usual address of each member at least five days prior to
the meeting.

 (3) Special meetings shall be called in the same manner by the
Chairman, upon the written request of three members of
the Panel.

(4) 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.

(5) All meetings of special committees will be called by their
respective chairmen pursuant to and in accordance with
performing their specified tasks.

(6) 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.

e. 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.

f. Avoidance of Conflicts of Interest

(1) Nongovernmental members of the Panel, and of special
committees established by the Panel, are "Special
Government Employees" within the meaning of NHB 1900.2A,
which sets forth guidance to NASA Special Government
employees regarding the avoidance of conflicts of interest
and the observance of ethical standards of conduct. A
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.

(2) Nongovernmental members of the Panel or a special 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.

6. SUPPORT

a. 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 the Panel.

b. 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) Insure that a written record is kept of all transactions, and
submit the same to the Panel for approval at each subsequent
meeting.

(3) Insure that the same service is provided for all special
committees of the Panel.

Administrator

CFR Title 14, Chapter 5, Subpart 1209.5.

     PART 1. 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 responsibili-
ties 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.

     The Board Chairman established a series of administrative procedures
to guide the Board's activities. In addition, specific assignments of
responsibility were made to all individuals involved in the Board's
activities so as to insure an efficient review activity. Overall logistic
and administrative support was provided by MSC.

     The Board conducted both Executive and General Sessions. During the
Executive Sessions, plans were agreed upon for guiding the Board's
activities and for establishing priorities for tests, analyses, studies,
and other Board efforts. At the General Sessions, status of Panel
activities was reviewed by the Board with a view towards coordination and
integration of all review activities. In addition, Board members regularly
attended daily status meetings of the Manned Spacecraft Center
Investigation Team.

     In general, the Board relied on Manned Spacecraft Center postmission
evaluation activities to provide the factual data upon which evaluation,
assessment, and analysis efforts could be based. However, the Board,
through a regular procedure, also levied specific data collection,
reduction, and analysis requirements on MSC. Test support for the Board
was conducted primarily at MSC but also included tests run at other NASA
Centers. Members of the Board and its Panels also visited a number of
contractor facilities to review manufacturing, assembly, and test
procedures applicable to the Apollo 13 mission.

     The Chairman of the Board provided the NASA Deputy Administrator with
oral progress reports. These reports summarized the status of Review Board
activities at the time and outlined the tasks still ahead. All material
used in these interim briefings was incorporated into the Board's official
files.

     As a means of formally transmitting its findings, determinations, and
recommendations, the Board chose the format of this Final Report which
includes both the Board's Judgments as well 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 MSC Investigation Team established a file of
data at MSC.

PART 2. BIOGRAPHIES OF BOARD MEMBERS, 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 10 years specialized in research on high-
speed aerodynamics there.

     In October 1958, Mr. Cortright was named Chief of Advanced Technology
Programs at NASA Headquarters, Washington, D. C., where he directed ini-
tial 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 Appli-
cations in 1965, 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 Astro-
nautics 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.

ROBERT F. ALLNUTT
NASA Headquarters

     Robert F, Allnutt, 34, Assistant to the NASA Administrator,
Washington, D C., is a member of the Apollo 13 Review Board.

     Mr. Allnutt was named to his present position this year. Prior to
that, he had been Assistant Administrator for Legislative Affairs since
1967.

     He joined NASA in 1960 as a patent attorney at the Langley Research
Center, Hampton, Virginia. In 1961, he was transferred to NASA
Headquarters Washington, D. C.

     Mr. Allnutt served as Patent Counsel for Communications Satellite
Corporation from January to September 1965, when he returned to NASA
Headquarters as Assistant General Counsel for Patent Matters.

     He is admitted to the practice of law in the District of Columbia and
the state of Virginia and is a member of the American Bar Association and
the Federal Bar Association.

     Mr. Allnutt was graduated from Virginia Polytechnic Institute with
a B.S. degree in industrial engineering. He received Juris Doctor and
Master of Laws degrees from George Washington University Law
School.

     Mr. and Mrs. Allnutt are the parents of two sons. The family lives
in Washington, D. C

NEIL A. ARMSTRONG
NASA Astronaut

     Neil A. Armstrong, 39, NASA astronaut, is a member of the Apollo 13
Review Board.

     Commander of the Apollo 11 mission and the first man on the Moon Mr.
Armstrong has distinguished himself as an astronaut and as an engineering
test pilot.

     Prior to joining the astronaut team at the Manned Spacecraft Center,
Houston, Texas, in 1962, Mr. Armstrong was an X-15 rocket aircraft project
pilot at the NASA Flight Research Center, Edwards, California.

     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 mal-
functioning 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 re-
ceived the Institute of Aerospace Sciences Octave Chanute Award, the AIAA
Astronautics Award, the NASA Exceptional Service Medal, the John F.
Montgomery Award, and the Presidential Medal of Freedom.

     He is a native of Wapakoneta, Ohio, and received a B.S. degree in
aeronautical engineering from Purdue University and a M.S. degree from the
University of Southern California. He was a naval aviator from 1949 to
1952 and flew 78 combat missions during the Korean action.

Mr. and Mrs. Armstrong have two sons.

JOHN F CLARK

NASA Goddard Space Flight Center

     Dr. John F. Clark, 49, Director of the NASA Goddard Space Flight
Center, Greenbelt, Maryland, is a member of the Apollo 13 Review
Board.

     He is an internationally known authority on atmospheric and space
sciences, holds four patents in electronic circuits and systems, and has
written many scientific papers on atmospheric physics, electronics, and
mathematics.

     Dr. Clark joined NASA in 1958 and served in the Office of Space
Flight Programs at NASA Headquarters until 1961 when he was named Director
of Geophysics and Astronomy Programs, Office of Space Sciences. From 1962
until 1965, he was Director of Sciences and Chairman of the Space Science
Steering Committee, Office of Space Science and Applications.

     In 1965, Dr. Clark was appointed Deputy Associate Administrator for
Space Science and Applications (Sciences), and later that year, Acting
Director of Goddard. He was named director of the center in 1966.

     Dr. Clark began his career in 1942 as an electronics engineer at the
Naval Research Laboratory, Washington, D.C. From 1947 to 1948 he was
Assistant Professor of Electronic Engineering at Lehigh University,
Bethlehem, Pennsylvania. He returned to NRL in 1948; and prior to Join-
ing 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., 48, 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 14th Air
Force, and as a project officer under Air Force Secretary Stuart
Symington. From 1952 to 1955, he was assigned to the Air Force Office of
Atomic 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 4951st 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 4925th Test Group.

General Hedrick Joined the Special Systems Office, Air Force Ballistics
Division, Los Angeles, in 1960. He was named Commander of 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 Techno-
logical 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.

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 1946, when he
was employed by the Argonne National Laboratory. In 1950, he joined the
AEC's Chicago Operations Office as staff chemical engineer. Later, he was
promoted to Assistant Manager for Technical Operations.  Generally engaged
in reactor development work for stationary power plants, he had a primary
role in the power reactor demonstration program.

     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 a research physicist at the
University's Lawrence Radiation Laboratory, Livermore, California, 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 Educa-
tion and the American Nuclear Society.

     Dr. Mark was born in Mannheim, Germany, and came to the United
States when he was 11 years old. He became a naturalized U.S. citizen
in 1945.

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

COUNSEL TO THE 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.

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 space-
craft 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

APOLLO 13 REVIEW BOARD OBSERVERS
WILLIAM A. ANDERS
National Aeronautics and Space Council

      William A. Anders, 36, Executive Secretary, National Aeronautics
and Space Council, Washington, D.C., is an official observer of the
Apollo 13 Review Board.

      Prior to being appointed to his present position in 1969, Mr.  Anders
was a NASA astronaut and an Air Force lieutenant colonel. He was lunar
module pilot on the Apollo & lunar orbital mission, man's first visit to
the vicinity of another celestial body.

      Mr. Anders joined the NASA astronaut team at the Manned Spacecraft
Center, Houston, Texas, in 1963. In addition to his Apollo 8 flight, he
served as backup pilot for Gemini 11 and backup command module pilot for
Apollo 11, the first lunar landing mission.

     Mr. Anders was commissioned a second lieutenant in the Air Force upon
graduation from the U.S. Naval Academy. After flight training, he served
as a pilot in all-weather interceptor squadrons of the Air Defense
Command. Prior to becoming an astronaut, he was a nuclear engineer and
instructor pilot at the Air Force Weapons Laboratory, Kirtland Air Force
Base, New Mexico.

     He is a member of the American Nuclear Society and has been awarded
the Air Force Commendation Medal, Air Force Astronaut Wings, the NASA
Distinguished Service Medal, and the New York State Medal for Valor.

     Mr. Anders was born in Hong Kong. He received a B.S. degree from the
U.S. Naval Academy and an M.S. degree in nuclear engineering from the Air
Force Institute of Technology.

Mr. and Mrs. 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.

     From 1941 to 1961, he was employed by the Mallinckrodt Chemical
Works, St. Louis, Missouri. Dr. Harrington started with the company as a
research chemist and in 1960, after a procession of research and
management positions, was appointed Vice President, Mallinckrodt Nuclear
Corporation and Vice President, Mallinckrodt Chemical Works.

     In 1961, when the fuel material processing plant of Mallinckrodt
became the Chemicals Division of United Nuclear Corporation, Dr.
Harrington was named Vice President of that division.

     He became Senior Vice President, United Nuclear Corporation,
Centreville, Maryland, in 1963.

     In 1965, Dr. Harrington was appointed President and General Manager,
Douglas United Nuclear, Inc. The company manages production reactors and
fuels fabrication facilities at Hanford, Washington, for the Atomic Energy
Commission.

     He is the co-author of a book, "Uranium 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. IRVING PINKEL
NASA Lewis Research Center

     I. Irving Pinkel, 57, Director, Aerospace Safety Research and Data
Institute at the NASA Lewis Research Center, Cleveland, Ohio, is an
official observer of the Apollo 13 Review Board.

     Until recently, he directed research at Lewis Research Center on
rocket propellant and electric power generation systems for space
vehicles, compressors and turbines for advanced aircraft engines, and
lubrication systems for rotating machines for these systems.

     Mr. Pinkel entered Government scientific service in 1935 as a
physicist with the U.S. Bureau of Mines, Pittsburgh, Pennsylvania. In
1940, he joined the staff of the Langley Research Center, Hampton,
Virginia, as a physicist. When the Lewis Research Center was built in
1942, he transferred there.

     He has been elected to Phi Beta Kappa, Sigma Xi, honorary scientific
society, and Pi Mu Epsilon, honorary mathematics fraternity. He is an Ohio
Professional Engineer, served on the former NACA subcommittees on
Meteorological Problems, Icing Problems, Aircraft Fire Prevention and
Flight Safety, and is a member of the NASA Research and Technology Advi-
sory Subcommittee on Aircraft Operating Problems. He has been a Special
Lecturer, Case Institute of Technology Graduate School.

     Mr. Pinkel has received the Flight Safety Foundation Award for con-
tributions to the safe utilization of aircraft, the Laura Taber Barbour
Award for development of a system for suppressing aircraft crash fires,
the NACA Distinguished Service Medal, and the NASA Sustained Superior
Performance Award.

     He was born in Gloversville, New York, and was graduated from the
University of Pennsylvania.

     Mr. and Mrs. Pinkel live in Fairview Park, Ohio. They are the
parents of two sons.

JAMES E. WILSON, JR.

Committee on Science and Astronautics
United States House of Representatives

     James E. Wilson, Jr., 39, Technical Consultant, United States House
of Representatives Committee on Science and Astronautics, is an official
observer of the Apollo 15 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 Head from 1956 to 1961.

     From 1954 to 1956, Mr. Wilson served as an officer in the U.S.  Army
Signal Corps. He was a development engineer with E. I. DuPont, Wilmington,
Delaware, from 1953 to 1954.

     Mr. Wilson is a member of Phi Sigma Alpha, a National Honor Society;
American Institute of Chemical Engineers; American Chemical Society, and
American Ordnance Association.

     Mr. Wilson is co-author of several publications of the House Commit-
tee on Science and Astronautics.

     He received a B.S. degree in chemical engineering from the Univer-
sity of Maine and a Master of Engineering Administration degree from
George Washington University.

     Mr. and Mrs. Wilson live in LaPlata, Maryland. They have two
children.

APOLLO 13 REVIEW BOARD PANEL CHAIRMEN

SEYMOUR C. HIMMEL
NASA Lewis Research Center

     Dr. Seymour C. Himmel, Assistant Director for Rockets and Vehicles,
Lewis Research Center, Cleveland, Ohio, heads the Design Panel of the
Apollo 13 Review Board.

     Dr. Himmel joined Lewis in 1948 as an aeronautical research scien-
tist. He has occupied supervisory positions since 1953.

     He has been awarded the NASA Exceptional Service Medal and the NASA
Group Achievement Award as manager of the Agena Project Group.  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. Himmel has a Bachelor of Mechanical Engineering degree from the
College of the City of New York and M.S. and Ph. D. degrees from Case
Institute of Technology.

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

EDWIN C. KILGORE
NASA Langley Research Center

     Edwin C. Kilgore, 47, Deputy Chief, Engineering and Technical Serv-
ices, Langley Research Center, Hampton, Virginia, heads the Project
Management Panel of the Apollo 13 Review Board.

     Mr. Kilgore joined the Langley science staff in 1944 and served in a
variety of technical and management positions until promotion to his
present position in 1968.

     He has received the Honorary Group Achievement Award for his role in
achieving a record of 97 consecutive successes for solid propellant rocket
motors and the NASA-Lunar Orbiter Project Group Achievement Award for
outstanding performance. He is a member of Pi Tau Sigma, honorary
mechanical engineering society.

     Mr. Kilgore was born in Coeburn, Virginia. He was graduated from
Virginia Polytechnic Institute with a B.S. degree in mechanical engi-
neering.

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 Lab-
oratory, 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 Achieve-
ment and Exceptional Service. In addition, he has received the Astro-
nautics 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, elec-
tronic tracking systems, and missile and range instrumentation.

     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.

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 recommenda-
tions by the Board could be based. The Panels were staffed with in-
dividual 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.

GENERAL ASSIGNMENTS FOR BOARD PANELS
(AS DOCUMENTED IN THE BOARD'S ADMINISTRATIVE PROCEDURES)

Panel 1 - Mission Events Panel

     It shall be the task of the Mission Events Panel to provide a de-
tailed 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 contin-
gency 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 be-
ginning 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 termina-
tion (Splash). Emphasis will be placed on the rationale used on mission
completion strategy.

Panel 1 Membership

     Mr. F. B. Smith, Panel Chairman
     Assistant Administrator for University Affairs
     NASA Headquarters
     Washington, D. C.

     Dr. Tom B. Ballard
     Aerospace Technologist
     Flight Instrument Division
     Langley Research Center
     Hampton, Virginia

     Mr. M. P. Frank
     Flight Director
     Flight Control Division
     Manned Spacecraft Center
     Houston, Texas

     Mr. John J. Williams
     Director, Spacecraft Operations
     Kennedy Space Center
     Florida

     Mr. Neil Armstrong, Board Member and Panel Monitor
     Astronaut
     Manned Spacecraft Center
     Houston, Texas

Panel 2 - Manufacturing and Test Panel

     The Manufacturing sad 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 pre-
flight test and checkout program, and any possible correlation of these
activities with the inflight events.

The Panel shall consist of three activities:

     Fabrication and Acceptance Testing - This will consist of reviewing
the fabrication, assembly, and acceptance testing steps actually used
during the manufacturing of the specific flight hardware elements in-
volved. 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 component-level
acceptance testing up through the countdown to lift-off for the specific
hardware involved. Test procedures and results will be reviewed as well as
observing specific tests where appropriate. Results of tests on other
serial number units will also be reviewed when appropriate.

     Reliability and Quality Assurance - This will be an overview of both
the manufacturing and testing, covering such things as parts and material
qualification and control, assembly and testing procedures, and inspection
and problem/failure reporting and closeout.

Panel 2 Membership

Mr. Harris M. Schurmeier, Panel Chairman
Deputy Assistant Laboratory Director for Flight Projects
Jet Propulsion Laboratory
Pasadena. California

Mr. Edward F. Baehr
Assistant Chief, Launch Vehicles Division
Deputy Manager, Titan Project
Lewis Research Center
Cleveland, Ohio

Mr. Karl L. Heimburg
Director, Astronautics Laboratory
Marshall Space Flight Center
Huntsville, Alabama

Mr. Brooks T. Morris
Manager, Quality Assurance and Reliability Office
Jet Propulsion Laboratory
Pasadena, California

Dr. John F. Clark, Board Member and Panel Monitor
Director
Goddard Space Flight Center
Greenbelt, Maryland

Panel 3 - Design Panel

     The Design Panel shall examine the design of the oxygen and asso-
ciated 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 re-
design. In addition, the Panel shall establish requirements for review of
other Apollo spacecraft systems of similar design.

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 com-
ponents, their derivation, changes thereto and the reasons therefore, and
the design of the system in response to the requirements, including such
elements as design approach, material selection, stress analysis, de-
velopment 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 manner
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
modifications.

     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 modifica-
tions 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. Browns Jr.
Chief, Strength of Materials Branch
Materials and Structures Division
Administration Directorate
Lewis Research Center
Cleveland, Ohio

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 Headquarters
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:

     1. Review and assess the effectiveness of the management struc-
ture employed in Apollo 13 in all areas pertinent to the Apollo 13
incident. This review will encompass the organization, the responsi-
bilities 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.

3. Review the project management lessons learned from the Apollo
13 mission from the standpoint of their applicability to subsequent
Apollo missions.

     Tasks 1 and 2, above, should encompass both the general review of the
processes used in Apollo 13 and specific applicability to the possible
cause or causes of the mission incident as identified by the Board.

Panel 4 Membership

E. C. Kilgore, Panel Chairman
Deputy Chief, Office of Engineering and Technical Services
Langley Research Center
Hampton, Virginia

R. D. Ginter
Director of Special Programs Office
Office of Advanced Research and Technology
NASA Headquarters
Washington, D.C.

Merrill H. Mead
Chief of Programs and Resources Office
Ames Research Center
Moffett Field, California

James B. Whitten
Assistant Chief, Aeronautical and Space Mechanics Division
Langley Research Center
Hampton, Virginia

Milton Klein, Board Member and Panel Monitor
Manager, AEC-NASA Space Nuclear Propulsion Office
Washington, D.C.

Board Observers

William A. Anders
Executive Secretary
National Aeronautics and Space Council
Washington, D.C.

Dr. Charles D. Harrington
Chairman
NASA Aerospace Safety Advisory Panel
Washington, D.C.

I. Irving Pinkel
Director
Aerospace Safety Research and Data Institute
Lewis Research Center
Cleveland, Ohio

Mr. James E. Wilson
Technical Consultant to the Committee on Science and Astronautics
United States House of Representatives
Washington, D.C.

Apollo 13 Review Board Support Staff

Brian M. Duff
Public Affairs Officer
Manned Spacecraft Center
Houston, Texas

Gerald J. Mossinghoff
Director of Congressional Liaison
NASA Headquarters
Washington, D.C.

Edward F. Parry
Counsel to Office of Manned Space Flight
NASA Headquarters
Washington, D.C.

Raymond G. Romatowski
Deputy Assistant Director for Administration
Langley Research Center
Hampton, Virginia

Ernest P. Swieda
Deputy Chief, Skylab Program Control Office
Kennedy Space Center, Florida

Consultants to the Board

Dr. Wayne D. Erickson, Head
Aerothermochemistry Branch
Langley Research Center
Hampton, Virginia

Dr. Robert Van Dolah
Acting Research Director
Safety Research Center
Bureau of Mines
Pittsburgh, Pennsylvania

MSC Support to the Board

     These persons were detailed by MSC to support the Apollo 13 Review
Board during its review activity at MSC. They are identified by MSC
Position title.

Roy C. Aldridge
Assistant to the Director of Administration

Mary Chandler
Secretary

Rex Cline
Technical Writer/Editor

Evon Collins
Program Analyst

Leroy Cotton
Equipment Specialist

Maureen Cruz
Travel Clerk

Janet Harris
Clerk Stenographer

Marjorie Harrison
Secretary

Phyllis Hayes
Secretary

William N. Henderson
Management Analyst

Sharon Laws
Secretary

Carolyn Lisenbee
Secretary

Judy Miller
Secretary

Jamie Moon
Technical Editor

Dorothy Newberry
Administrative Assistant

Lettie Reed
Editorial Assistant

Charlene Rogozinski
Secretary

Joanne Sanchez
Secretary

Billie Schmidt
Employee Development Specialist

Frances Smith
Secretary

George Sowers

Management Presentations Officer

Elaine Stemerick
Secretary

2-26
Mary Thompson
Administrative Assistant

Alvin C. Zuehlke
Electrical Engineer

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 Adminis-
trator, and key NASA officials in Washington, D.C., to discuss Board
membership.

     The Chairman met with NASA Office of Manned Space Flight top offi-
cials while enroute to MSC on NASA aircraft and discussed program organi-
zation 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.

APRIL 22, 1970

     The Board met with Colonel McDivittts MSC Investigation Team to re-
view the progress made by MSC in identifying causes of the accident and in
developing an understanding of sequences and relationships between known
inflight events. In addition, MSC officials briefed the Board on MSC
Investigation Team structure and assignments.

     The Board met with Panel 1 of the MSC Investigation Team for de-
tailed discussion of inflight events and consideration of early con-
clusions on implications of preliminary data analysis.

     The Board held its second meeting to discuss MSC investigative
efforts and additional appointments of Panel specialists.

     Board members attended Panel 1 evening roundup of day's evaluation
activities, which included detailed discussions of specific studies, data
reductions, and support test activities already underway.

APRIL 23, 1970

     The Apollo 15 Review Board established itself in proximity to the MSC
Investigation Team in 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 made to 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, par-
ticularly to provide efficient interface with MSC personnel.

     Board and Panel Members again met with MSC officials to further re-
view the sequence of events in the Apollo 15 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 NASA specialists in order to
bring Panels up to strength.

APRIL 24, 1970

     Board Members, Panel Chairmen, and MSC officials reviewed additional
data analysis made by MSC and contractor personnel with particular empha-
sis 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 loca-
tions.

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 cryo-
genic oxygen tanks.

APRIL 25, 1970

     The Board met to discuss details of onsite inspections of command
service module (CSM) flight hardware at principal contractor installa-
tions.

     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
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, 1970

     Chairman Cortright outlined a plan for the Board's preliminary report
scheduled for presentation to the Deputy Administrator during his visit to
MSC on May 1. Each Panel Chairman was to summarize the status of his
Panel's activities for Dr. George Low on Friday, April 29, 1970.

     Board Member Neil Armstrong completed arrangements to provide each
Board Member and Panel Chairman an opportunity for detailed simulation of
the Apollo 13 inflight accident using MSC's CSM simulation equipment.

     Board and Panel Members reviewed enhanced photographs of the
Apollo 13 service module at the MSC Photographic Laboratory.

     Dr. von Elbe of Atlantic Research Company briefed Board and Panel
Members on cryogenics and combustion phenomena.

     A representative of the Manufacturing and Test Panel performed an
onsite inspection at Beech Aircraft, Boulder.

     Manufacture and Test Panel-personnel reviewed detanking procedures
followed at KSC during the Apollo 13 countdown demonstration test (CDDT).

     Board and Panel personnel reviewed progress to date at a general
Board meeting involving all Review Board personnel.

APRIL 29, 1970

     Dr. Charles Harrington, Board Observer and Chairman of the Aerospace
Safety Advisory Panel, arrived for a 2-day detailed review of Board pro-
cedures and Progress in the accident review.

     The Board reviewed North American Rockwell preliminary recommenda-
tions involving oxygen tank redesign.

     The Board continued to review and examine oxygen tank ignition
sources and combustion propagation processes with specialists from MSC,
other NASA Centers, 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 pro-
gress to date and to discuss work plans for the next several days.

     The Project Management Panel visited North American Rockwell at
Downey to review detailed procedures for acceptance tests, subcontractor
inspections, project documentation, and other management interface areas.

APRIL 30, 1970

     The Safety Advisory Panel continued discussions with Board Chairman
and MSC officials on progress of total Apollo 13 review efforts.

     Panel Members reviewed instrumentation used in Apollo 13 spacecraft
in order to establish the validity of telemetry data being used in Board
analysis.

     Chairman Cortright convened two Board meetings to review progress of
the work and to discuss preliminary findings of the Board.

     Project Management personnel visited Beech Aircraft Corporation to
review procedures used for assembly of cryogenic oxygen tanks and to dis-
cuss communication and information systems within the APOLLO Program.

Panels continued to review detailed data in their respective areas.

MAY 1, 1970

     Board and Panel personnel participated in a joint MSC/Apollo 13
Review Board status presentation to the NASA Deputy Administrator.  The
meeting covered all significant Apollo 13 findings and early conclusions
on the cause of the accident and appropriate remedial actions.

     The MSC staff briefed Board Members on initiaI evaluations of pro-
posed design changes in oxygen tank system.

     Panel Members continued to assess data accumulated from the Apollo 13
mission with particular emphasis upon the design and performance of elec-
tric power systems used in the service module.

     Board Members and Panel Chairmen reviewed specific test matrix being
proposed by Apollo 13 Review Board specialists covering most significant
unknowns involved in understanding failure mechanisms.

MAY 2, 1970

     Board Members met in General Session to discuss preparation of a com-
plete "failure tree" as an additional guide in conducting a complete re-
view and investigation. Specific aspects of this approach were reviewed.

     The Project Management Panel reviewed oxygen tank reliability history
and quality assurance criteria used in assembly, test, and checkout of
these systems.

     Panel specialists continued reviewing data from the mission with
emphasis upon integrating various data points into logical failure mode
patterns established by MSC and Board personnel.

MAY 3 1970

     Chairman Cortright and Board Members conducted a detailed review of
individual Panel status and progress and established milestones for
additional analytical work and preparation of preliminary findings.

     The Board and Panel agreed to tentative report structure, including
required exhibits, tables, drawings, and other reference data.

     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 Members worked on individual analyses with particular attention
to developing requirements for additional test activity in support of ten-
tative conclusions.

     The Board agreed to strengthen its technical reviews of combustion
propagation and electrical design by adding specialists in these areas.

MAY 4, 1970

     The Design Panel continued its intensive review of the "shelf drop"
incident at NR involving 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 review activity.

The Board continued to meet with individual Panels and support
offices to review the status of preliminary findings and work
completed.

MAY 5, 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.

     MSC personnel briefed Panel Members on availability of additional
telemetry data in the MSC data bank in order to insure Board considera-
tion of all possible useful data.

     Panels commenced initial drafting of preliminary findings in specific
areas, including summary descriptions of system performance during the
Apollo 13 flight.

     The Board met with the MSC Investigation Team for complete review of
the proposed test program.

MAY 6, 1970

     Board Members, MSC personnel, and Members of NASA's Aerospace Safety
Advisory Panel met for detailed discussions and evaluation of accident
review status and progress. The review covered oxygen tank questions,
recovery operations, and a mission simulation by MSC astronauts.

     Panel Members continued to work on the preparation of preliminary
Panel drafts.

     Chairman Cortright transmitted additional requests for tests to MSC
and modified procedures for control of overall test activity relating to
the Apollo 13 accident.

MAY 7, 1970

     The General Board Session reviewed complete analysis and test support
activities being conducted for the Board and MSC at various governmental
and contractor installations.

     Board and Panel Members met to discuss Ames laboratory tests con-
cerning liquid oxygen combustion initiation energies required in the
cryogenic oxygen tank used in the Apollo 13 SM.

     Panel 1 Members reviewed mission control equipment and operating
procedures used during the Apollo 13 mission and reviewed actual mission
events in detail.

     The Panels continued to develop preliminary drafts of their reviews
and analyses for consideration by the Board.

MAY 8, 1970

     Dr. Robert Van Dolah, Bureau of Mines, joined the Board as a con-
sultant on combustion propagation and reviewed Apollo 13 Review Board data
developed to date.

     The General Board Session convened to review proposed report format
and scope. An agreement was reached on appendices, on the structure of the
report, and on the degree of detail to be included in individual Panel
reports.

     Chairman Cortright assigned additional specific test overview re-
sponsibilities to members of the Apollo 13 Review activity.

     Panel 1 conducted a formal-interview with the MSC Flight Director
covering all significant mission events from the standpoint of ground
controllers.

     Panels 2 through 4 continued developing preliminary reports.  Panel 4
announced a formal schedule of interviews of MSC, contractors, and NASA
Headquarters personnel.

     Board Members explored in detail possible failure mode sequences
developed by MSC personnel involving ignition and combustion within the SM
cryogenic oxygen tank.

     The Board recessed for 3 days, leaving a cadre of personnel at MSC to
edit preliminary drafts developed by the Panels and to schedule further
activity for the week of May 11.

MAY 9, 1970

Board in recess.

MAY 10, 1970

Board in recess.

MAY 11, 1970

     Board in recess. MSC support personnel continued work obtaining
additional technical data for Board review.

MAY 12, 1970

Board Members returned to MSC.

     Board Members attended a General Session to review progress and
status of the report.

     Panel Chairmen reported on individual progress of work and estab-
lished schedules for completion of analyses and evaluations.

     Chairman Cortright reported on the Langley Research Center support
test program aimed at simulation of SM panel ejection energy pulses.

MAY 13, 1970

     Board Members reviewed preliminary drafts of report chapter on Re-
view and Analysis and Panel 1 report on Mission Events.

     Mission Events Panel Members interviewed Electrical, Electronic, and
Communications Engineer (EECOM) and one of the Apollo 13 Flight Directors
on activities which took place in the Mission Control Center (MCC) during
and after the flight accident period.

     Panel 4, Project Management Panel, conducted interviews with princi-
pal Apollo 13 program personnel from MSC and contract organizations.

     Panel Members continued drafting preliminary versions of Panel re-
ports for review by the Board.

     Manufacturing and Test Panel representatives discussed program for
oxygen tank testing to be conducted at Beech Aircraft.

     Board Members met in General Session to review report milestones and
required test data for the week ahead.

MAY 14, 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 Management Panel continued to interview key Apollo project
personnel from NASA Centers and contractors.

     Panel Members circulated first drafts of all Panel reports to Board
Members for review and correction.

MAY 15, 1970

     Mission Events Panel personnel interviewed Apollo 13 Command Module
Pilot John Swigert to verify event chronology compiled by the Panel and to
review crew responses during Apollo 13 mission.

     Project Management Panel continued interviewing key project personnel
with NASA Centers and contractors.

     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.

     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 Beech Aircraft.

MAY 19. 1970

     Board Members met in Executive Session to continue evaluation and
assessment of preliminary findings, determinations, and recommendations
prepared by individual Board Members and 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 with Mr. 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.

     Board Members and others met with MSC officials 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 recommen-
dations was developed and circulated by the Chairman for review and
comment.

     Arrangements were made with NASA Headquarters officials for pack-
aging, delivery, and distribution of the Board's final report.

     Mission Events Panel conducted an interview with Lunar Module Pilot
IIaise to review selected mission events bearing on the accident.

MAY 22, 1970

     Mission Events Panel representatives met with MSC officials 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 con-
ducted in support of the Apollo 15 Review.

     Board Observer I. I. Pinkel briefed the Board on Lewis Research
Center fire propagation tests involving Teflon.

MAY 23, 1970

     Board Members reviewed Chapter 4 of Board report entitled "Review and
Analysis."

     Panel Chairmen reviewed draft findings and determinations prepared by
the Board.

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 Session 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.

     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.

MAY 28, 1970

Board in recess.

MAY 29, 1970

Board in recess.

MAY 30, 1970

Board in recess.

MAY 31, 1970

Board in recess.

JUNE 1, 1970

Board Members returned to MSC.

     Board and Panel Members met in General Session to discuss revisions
of Panel reports in light of latest information regarding thermostatic
switch failure during CDDT at KSC.

     Board approved new schedule for Board report calling for final
versions of Panel reports by Monday, June 8.

JUNE 2, 1970

     Chairman Cortright briefed the Press on the status of the Board's
work and future plans.

     Board and Panel Members participated in a detailed interview and
discussion with MSC and contractor personnel regarding specific coordina-
tion steps taken during oxygen tank no. 2 detanking operations at KSC.

     Board Members met in Executive Session to review latest test results
and to assess status of Board findings and determinations.

JUNE 3, 1970

     Board and Panel Members met with MSC Program Office personnel for a
detailed update of recent MSC information and analyses stemming from on-
going test programs.

     Board Members and Panel Chairmen completed final reviews of Panel
reports and also reviewed final draft of findings, determinations, and
recommendations.

     Board and Panel Members received a detailed briefing on thermostatic
switch questions with emphasis upon actions of various organizations
during and after detanking operations at KSC.

JUNE 4, 1970

     Board Members met in Executive Session and completed final revisions
of Chapter 4 of the Board summary.

     Board and Panel Members witnessed a special full-scale tank ignition
test performed at MSC.

     Panel Chairmen completed final revisions of individual Panel reports
and submitted copy to the Reports Editorial Office.

     Board met in Executive Session and agreed to final schedule for re-
port printing and delivery to the Administrator on June 15, 1970.

JUNE 5, 1970

     Board Members met in Executive Session and completed work on Chap-
ter 5 of the Board Summary Report (Findings, Determinations, and Recom-
mendations).

     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.

CHAPTER 3
DESCRIPTION OF APOLLO 13 SPACE VEHICLE
AND MISSION SUMMARY

     This chapter is extracted from Mission Operation Report No. M-932-70,
Revision 3, published by the Program and Special Reports Division (XP),
Executive Secretariat, NASA Headquarters, Washington, D.C.

     Discussion in this chapter is broken into two parts. Part l 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.

PART 1 APOLLO/SATURN V SPACE VEHICLE

     The primary flight hardware of the Apollo Program consists of the
Saturn V launch vehicle and Apollo spacecraft (fig. 3-1). 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-1 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 electri-
cally with the S-II stage. It also interfaces structurally, elec-
trically, 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 inde-
pendent 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.

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 h.5:1) during the first 100
seconds of translunar injection. The remaining translunar injection
acceleration is provided at a thrust level of 203,000 pounds (mixture mass
ratio of 5.0:1). 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).

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.  Aerodynami-
cally, the SLA smoothly encloses the irregularly shaped LM and transitions
the space vehicle diameter from that of the upper stage of the LV to that
of the SM. The SLA also encloses the nozzle of the SM engine and the high
gain antenna.

     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 pro-
pulsion 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
an 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 section 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.

     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 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 (UDMX) 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 SM are
the steerable S-band high-gain antenna, mounted on the aft bulkhead; two
VHF omnidirectional antennas, mounted on opposite sides of the module near
the top; and the rendezvous radar transponder antenna, mounted in the SM
fairing.

     Seven lights are mounted in the aluminum panels of the fairing.
Four lights (one red, one green, and two amber) are used to aid the
astronauts in docking: one is a floodlight which can be turned on to
give astronauts visibility during extravehicular activities, one is a
flashing beacon used to aid in rendezvous, and one is a spotlight used
in rendezvous from 500 feet to docking with the LM.

     SM/CM separation.- Separation of the SM from the CM occurs shortly
before entry. The sequence of events during separation is controlled
automatically by two redundant service module jettison controllers (SMUC)
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 steel- brazed
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 cotrols 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.

     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, environ-
mental control, and cryogenic storage systems. The rotation and trans-
lation controllers used for attitude, thrust vector, and translation
maneuvers are located on the arms of two crew couches. In addition, a
rotation controller can be mounted at the navigation position in the lower
equipment bay.

     Critical conditions of most spacecraft systems are monitored by a
caution and warning system. A malfunction or out-of-tolerance condition
results in illumination of a status light that identifies the abnormal-
ity. It also activates the master alarm circuit, which illuminates two
master alarm lights on the MDC and one in the lower equipment bay and
sends an alarm tone to the astronauts' headsets. The master alarm lights
and tone continue until a crewman resets the master alarm circuit. This
can be done before the crewmen deal with the problem indicated. The
caution and warning system also contains equipment to sense its own
malfunctions.

Lunar Module

     The lunar module (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.

     Main Propulsion- Main propulsion is provided by the descent pro-
pulsion 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 unsymmetrical- dimethylhydrazine 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.

The AGS commands on/off operation but has no automatic throttle control
capability. Manual control capability of engine firing functions has been
provided. Manual thrust control override may, at any time, command more
thrust than the level commanded by the LM guidance computer (LGC).

     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) con-
tains 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 115-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 capacity at 28 volts.
The ascent stage batteries are normally connected in parallel for even
discharge. Because of design load characteristics, the ascent stage
batteries do not have and do not require high and low voltage taps.

     Nominal voltage for ascent stage and descent stage batteries is 30.0
volts. Reverse current relays for battery failure are one of many
components designed into the FPS to enhance EPS reliability. Cooling of
the batteries is provided by the environmental control system cold rail
heat sinks. Available ascent electrical energy is 17.8 kilowatt hours-at a
maximum drain of 50 amps per battery and descent energy is 46.9 kilowatt
hours at a maximum drain of 25 amps per battery.

MISSION MONITORING, SUPPORT, AND CONTROL

     Mission execution involves the following functions: prelaunch
checkout and launch operations; tracking the space vehicle to determine
its present and future positions; securing information on the status of
the flight crew and space vehicle systems (via telemetry); evaluation of
telemetry information; commanding the space vehicle by transmitting
real-time and updata commands to the onboard computer; and voice com-
munication between flight and ground crews.

     These functions require the use of a facility to assemble and launch
the space vehicle (see Launch Complex), a central flight control facility,
a network of remote stations located strategically around the world, a
method of rapidly transmitting and receiving information between the space
vehicle and the central flight control facility, and a real-time data
display system in which the data are made available and presented in
usable form at essentially the same time that the data event occurred.

     The flight crew and the following organizations and facilities
participate in mission control operations:

     a. Mission Control Center (MCC), Manned Spacecraft Center (MSC),
Houston, Texas. The MCC contains the communication, computer display, and
command systems to enable the flight controllers to effectively monitor
and control the space vehicle.

     b. Kennedy Space Center (KSC), Cape Kennedy, Florida. The space
vehicle is launched from KSC and controlled from the Launch Control Center
(LCC). Prelaunch, launch, and powered flight data are collected at the
Central Instrumentation Facility (CIF) at KSC from the launch pads, CIF
receivers, Merritt Island Launch Area (MILA), and the down-range Air
Force Eastern Test Range (AFETR) stations. These data are transmitted to
MCC via the Apollo Launch Data System (ALDS). Also located at KSC (AFETR)
is the Impact Predictor (IP), for range safety purposes.

     c. Goddard Space Flight Center (GSFC), Greenbelt, Maryland. GSFC
manages and operates the Manned Space Flight Network (MSFN) and the NASA
communications (NASCOM) network. During flight, the MSFN is under the
operational control of the MCC.

     d. George C. Marshall Space Flight Center (MSFC), Huntsville,
Alabama. MSFC, by means of the Launch Information Exchange Facility (LIEF)
and the Huntsville Operations Support Center (HOSC) provides launch
vehicle systems real-time support to KSC and MCC for preflight, launch,
and flight operations.

     A block diagram of the basic night control interfaces is shown
in figure 3-10.

Vehicle Flight Control Capability

     Flight operations are controlled from the MCC. The MCC has two flight
control rooms, but only one control room is used per mission. Each control
room, called a Mission Operations Control Room (MOCR), is capable of
controlling individual Staff Support Rooms (SSR's) located adjacent to the
MOCR. The SSR's are manned by flight control specialists who provide
detailed support to the MOCR. Figure 3-11 outlines the organization of the
MCC for flight control and briefly describes key responsibilities.
Information flow within the MOCR is shown in figure 3-12.

     The consoles within the MOCR and SSR's permit the necessary inter-
face between the flight controllers and the spacecraft. The displays and
controls on these consoles and other group displays provide the capability
to monitor and evaluate data concerning the mission and, based on these
evaluations, to recommend or take appropriate action on matters concerning
the flight crew and spacecraft.

     Problems concerning crew safety and mission success are identified to
flight control personnel in the following ways:

a. Flight crew observations

b. Flight controller real-time observations

c. Review of telemetry data received from tape recorder playback

d. Trend analysis of actual and predicted values

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

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
(ALSEP).

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

Launch and Earth Parking Orbit

     Apollo 13 was successfully launched on schedule from Launch Complex
39A, Kennedy Space Center, Florida, at 2:13 p.m. e.s.t., April 11, 1970.
The launch vehicle stages inserted the S-IVB/instrument unit (IU)/
spacecraft combination into an earth parking orbit with an apogee of 100.2
nautical miles (n. mi.) and a perigee of 98.0 n. mi. (100-n.-mi. circular
planned). 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 34 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 3-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 made for translunar injection (TLI).
Onboard television was initiated at 01:35 ground elapsed time (g.e.t.) for
about 5.5 minutes. The second S-IVB burn was initiated on schedule for
TLI. All major systems operated satisfactorily and all end conditions
were nominal for a free-return circumlunar trajectory.

Translunar Coast

     The CSM separated from the LM/IU/S-IVB at about 03:07 g.e.t.  On-
board television was then initiated for about 72 minutes and clearly
showed CSM "hard docking,"-ejection of the CSM/LM from the S-IVB at about
04:01 g.e.t., and the S-IVB auxiliary propulsion system (APS) evasive
maneuver as well as spacecraft interior and exterior scenes. The SM RCS
propellant usage for the separation, transposition, docking, and ejection
was nominal. All launch vehicle safing activities were performed as
scheduled.

     The S-IVB APS evasive maneuver by an 8-second APS Ullage burn was
initiated at 04:18 g.e.t. and was successfully completed. The liquid
oxygen dump was initiated at 04:39 g.e.t. and was also successfully
accomplished. The first S-IVB ALPS burn 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
653' S., 3053' W. versus the targeted 3 S., 30 W. Therefore the second
S-IVB APS (trim) burn was not required. The gaseous nitrogen pressure
dropped in the IU ST-124-M3 inertial platform at 18:25 g.e.t.  and the
S-IVB/IU no longer had attitude control but began tumbling slowly.
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 l2 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 11:41 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 re-
ported an undervoltage alarm on the CSM main bus B. Pressure was rapid-
ly lost in 5M 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 mid-
Pacific Ocean at 142:53 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 LM was powered down to
conserve consumables; only the LM environmental control system (ECS) and
communications and telemetry systems were kept powered up.

     The LM DPS was used to perform MCC-5 at 105:19 g.e.t. The 15-second
burn (at 10-percent throttle) produced a velocity change of about 7.8 fps
and successfully raised the entry flight path angle to -6.52.

     The CSM was partially powered up for a check of the thermal condi-
tions 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
high-gain 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 10 miles; 6-knot ENE winds;
and wave height 1 to 2 feet. Drogue and main parachutes deployed normally.
Visual contact with the spacecraft was reported at 142:50 g.e.t.  Landing
occurred at 142:54:41 g.e.t. (01:07:41 p.m. e.s.t., April 17). The landing
point was in the mid-Pacific Ocean, approximately 2140' S., 16522' 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 heli-
copter, was safe aboard the ship at 1:53 p.m. e.s.t., less than an hour
after landing.

CHAPTER 4
REVIEW AND ANALYSIS OF APOLLO 13 ACCIDENT

PART 1. INTRODUCTION

     It became clear in the course of the Board's review that the acci-
dent during the Apollo 13 mission was initiated in the service module
cryogenic oxygen tank no. 2. Therefore, the following analysis centers on
that tank and its history. In addition, the recovery steps taken in the
period beginning with the accident and continuing to reentry are
discussed.

     Two oxygen tanks essentially identical to oxygen tank no. 2 on Apollo
13, and two hydrogen tanks of similar design, operated satisfactorily on
several unmanned Apollo flights and on the Apollo 7, 8, 9, 10, 11, and 12
manned missions. With this in mind, the Board placed particular emphasis
on each difference in the history of oxygen tank no. 2 from the history of
the earlier tanks, in addition to reviewing the design, assembly, and test
history.

PART 2. OXYGEN TANK NO. 2 HISTORY

DESIGN

     On February 26, 1966, the North American Aviation Corporation, now
North American Rockwell (RR), 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 be-
tween 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 connector 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 capaci-
tance 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
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 super-
critical 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 1000 psi. The oxygen tank dome is open to the
vacuum between the inner and outer tank shell and contains a rupture disc
designed to blow out at about 75 psi.

     The approximate amounts of principal materials within the oxygen
tank are set forth in table 4-I.

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 of the
Apollo 13 service module (SM 109) at the North American Rockwell plant
prior to shipment to KSC. Figure 4-5 shows the 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 shows the oxygen tank shelf, with oxygen tank no. 2 at left
center. Figure 4-7 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
some of the oxygen system instrumentation and wiring, largely covered by
insulation. Figure 4-8 is a photograph of the bay 4 panel, which was
missing from the service module after the accident.

     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 subcon-
tracts with Beech, the inner shell of the tank was manufactured by the
Airite Products Division of Electrada Corporation; the quantity probe was
made by Simmonds Precision Products, Inc., and the fans and fan motors
were produced by Globe Industries, Inc.

     The Beech serial number assigned to the oxygen tank no. 2 flown in
the Apollo 13 was 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 assembled and welded, the heater
assembly must be inserted in the tank, moved to one side, and bolted in
place. Then the quantity probe is inserted into the tank and the heater
assembly wires (to the heaters, the thermostats, and the fan motors) must
be pulled through the head of the quantity probe and the 32-inch coiled
conduit in the dome. Thus, the design requires during assembly a
substantial amount of wire movement inside the tank, where movement cannot
be readily observed, and where possible damage to wire insulation by
scraping or flexing cannot be easily detected before the tank is capped
off and welded closed.

     Several minor manufacturing flaws were discovered in oxygen tank no.
2 in the course of testing. A porosity in a weld on the lower half of the
outer shell necessitated grinding and rewelding. Rewelding was also
required when it was determined that incorrect welding wire had been
inadvertently used for a small weld on a vacuum pump mounted on
the outside of the tank dome. The upper fan motor originally installed was
noisy and drew excessive current. The tank was disassembled and the heater
assembly, fans, and heaters were replaced with a new assembly and new
fans. The tank was then assembled and sealed for the second time, and the
space between the inner and outer shells was pumped down over a 28-day
period to create the necessary vacuum.

TANK TESTS AT BEECH

     Acceptance testing of oxygen tank no. 2 at Beech included extensive
dielectric, insulation, and functional tests of heaters, fans, and vac-
uum 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 900 psi for 25
to 30 hours over a range of ambient conditions and out-flow 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 applying warm gas at about 30 psi to the vent line to force
the liquid oxygen (LOX) in the tank out the fill line (see fig. 4-2). No
difficulties were recorded in this detanking operation.

     The acceptance test indicated that the rate of heat leak into the
tank was higher than permitted by the specifications.  After some re-
working, 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 10024XTA0008 as oxygen tank no. 2, was completed on March 11,
1968. The shelf was to be installed in SM 106 for flight in the Apollo 10
mission.

      Beginning on April 27, the assembled oxygen shelf underwent stand-
ard proof-pressure, leak, and functional checks. One valve on the shelf
leaked and was repaired, but no anomalies were noted with regard to oxygen
tank no. 2, and therefore no rework of oxygen tank no. 2 was required.
None of the oxygen tank testing at NR requires use of LOX in the tanks.

     On June 4, 1968, the shelf was installed in SM 106.

     Between August 3 and August 8, 1968, testing of the shelf in the SM
was conducted. No anomalies were noted.

     Due to electromagnetic interference problems with the vac-ion pumps
on cryogenic tank domes in earlier Apollo spacecraft, a modification was
introduced and a decision was made to replace the complete oxygen shelf in
SM 106. 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 4. One shelf
bolt was mistakenly left in place during the initial attempt to remove the
shelf; and as a consequence, after the front of the shelf was raised about
2-inches, the fixture broke, allowing the shelf to drop back into place.
Photographs of the underside of the fuel cell shelf in SM 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 without 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. 1510
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
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
5mm Hg followed by an oxygen pressure of about 80 psi. After the cooling
of the fuel cells, cryogenic oxygen loading and tank pressurization to 331
psi were completed without abnormalities. At the time during CDDT when the
oxygen tanks are normally partially emptied to about 50 percent of
capacity, oxygen tank no. 1 behaved normally, but oxygen tank no. 2 only
went down to 92 percent of its capacity. The normal procedure during CDDT
to reduce the quantity in the tank is to apply gaseous oxygen at 80 psi
through the vent line and to open the fill line. When this procedure
failed, it was decided to proceed with the CDDT until completion and then
look at the oxygen detanking problem in detail. An Interim Discrepancy
Report was written and transferred to a Ground Support Equipment (GSE)
Discrepancy Report, since a GSE filter was suspected.

     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 that the problem might be due to a leak in the
path between the fill line and the quantity probe due to loose fit in the
sleeves and tube. Referring to figure 4-2, it will be noted that such a
leak would allow the gaseous oxygen (GOX) being supplied to the vent line
to leak directly to the fill line without forcing any significant amount
of LOX out of the tank. At this point, a discrepancy report against the
spacecraft system was written.

     A "normal" detanking procedure was then conducted on both oxygen
tanks, pressurizing through the vent line and opening the fill lines. Tank
no. 1 emptied in a few minutes. Tank no. 2 did not. Additional attempts
were made with higher pressures without effect, and a decision was made to
try to "boil off" the remaining oxygen in tank no. 2 by use of the tank
heaters. The heaters were energized with the 65 V dc. GSE power supply,
and, about 1-1/2 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, KSC personnel consulted with cognizant personnel at
MSC and at NR and decided to test whether the oxygen tank no. 2 could be
filled without problems. It was decided that if the tank could be filled,
the leak in the fill line would not be a problem in flight, since it was
felt that even a loose tube resulting in an electrical short between the
capacitance plates of the quantity gage would result in an energy level
too low to cause any other damage.

     Replacement of the oxygen shelf in the CM would have been difficult
and would have taken at least 45 hours. In addition, shelf replacement
would have had the potential of damaging or degrading other elements of
the SM in the course of replacement activity. Therefore, the decision was
made to test the ability to fill oxygen tank no. 2 on March 30,
1970, twelve days prior to the scheduled Saturday, April 11, launch, so as
to be in a position to decide on shelf replacement well before the launch
date.

     Accordingly, flow tests with GOX were run on oxygen tank no. 2 and on
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 showed
that very low energy levels would result. On the filling test, oxygen
tanks no. 1 and no. 2 were filled with LOX to about 20 percent of capacity
on March 30 with no difficulty. Tank no. 1 emptied in the normal manner,
but emptying oxygen tank no. 2 again required pressure cycling with the
heaters turned on.

     As the launch date approached, the oxygen tank no. 2 detanking
problem was considered by the Apollo organization. At this point, the
"shelf drop" incident on October 21, 1968, at NR was not considered and it
was felt that the apparently normal detanking which had occurred in 1967
at Beech was not pertinent because it was believed that a different
procedure was used by Beech. In fact, however, the last portion of the
procedure was quite similar, although a slightly lower GOX pressure was
utilized.

     Throughout these considerations, which involved technical and
management personnel of KSC, MSC, NR, Beech, and NASA Headquarters,
emphasis was directed toward the possibility and consequences of a loose
fill tube, very little attention was paid to the extended operation of
heaters and fans except to note that they apparently operated during and
after the detanking sequences.

     Many of the principals in the discussions were not aware of the
extended heater operations. Those that did know the details of the
procedure did not consider the possibility of damage due to excessive heat
within the tanks and therefore did not advise management officials of any
possible consequences of the unusually long heater operations.

     As noted earlier in this chapter, and shown in figure 4-2, each
heater is protected with a thermostatic switch, mounted on the heater
tube, which is intended to open the heater circuit when it senses a
temperature of 80 F. In tests conducted at MSC since the accident,
however, it was found that the switches failed to open when the heaters
were powered from a 65 V dc supply similar to the power used at KSC during
the detanking sequence. Subsequent investigations have shown that the
thermostatic switches used, while rated as satisfactory for the 28 V dc
spacecraft power supply, could not open properly at 65 V dc. Qualification
and test procedures for the heater assemblies and switches do not at any
time test the capability of the switches to open while under full current
conditions. A review of the voltage recordings made during the detanking
at KSC indicates that, in fact, the switches did not open when the
temperature indication from within the tank rose past 80 F.  Further
tests have shown that the temperatures on the heater tube may have
reached as much as 1000 F during the detanking. This temperature will
cause serious damage to adjacent Teflon insulation, and such damage almost
certainly occurred.

     None of the above, however, was known at the time and, after
extensive consideration was given to all possibilities of damage from
a loose fill tube, it was decided to leave the oxygen shelf and oxygen
tank no. 2 in the SM and to proceed with preparations for the launch
of Apollo 13.

     The manufacture and test history of oxygen tank no. 2 is discussed
in more detail in Appendix C to this report.

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 11, 1970, from the
KSC Launch Complex 39A. The spacecraft was inserted into a 100-nautical-
mile 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 plan-
ned. 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.

     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 Launch Area,
Florida, tracking station.

     The S-IVB restarted at 02:35:46 g.e.t for the translunar injection
burn, with shutdown coming some 5 minutes 51 seconds later.  Accuracy of
the Saturn V instrument unit guidance for the TLI burn was such that a
planned midcourse correction maneuver at 11:41:23 g.e.t. was not neces-
sary.  After TLI, Apollo 13 was calculated to be on a free-return trajec-
tory with a predicted closest approach to the lunar surface of 210
nautical miles.

     The CSM was separated from the S-IVB about 3 hours after launch, and
after a brief period of stationkeeping, the crew maneuvered the CSM to
dock with the LM vehicle in the LM adapter atop the S-IVB stage.  The
S-IVB stage was separated from the docked CSM and LM shortly after 4 hours
into the mission.

     In manned lunar missions prior to Apollo 13, the spend S-IVB third
stages were accelerated into solar orbit by a "slingshot" maneuver in
which residual liquid oxygen was dumped through the J-2 engine to pro-
vide propulsive energy.  On Apollo 13, the plan was to impact the S-IVB
stage on the lunar surface in proximity to the seismometer emplaced in the
Ocean of Storms by the crew of Apollo 12.

     Two hours after TLI, the S-IVB attitude thrusters were ground com-
manded on to adjust the stage's trajectory toward the designated impact at
latitude 3 degrees S. by longitude 30 degrees W.  Actual impact was at
latitude 2.4 degrees S. by longitude 27.9 degrees 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 Moon lasted almost 4 hours
and provided lunar scientists with additional data on the structure of the
Moon.

As in previous lunar missions, the Apollo 13 spacecraft was set up in the
passive thermal control (PTC) mode which calls for a continuous roll rate
of three longitudinal axis revolutions each hour.  During crew rest
periods and at other times in translunar and transearth coast when a
stable attitude is not required, the spacecraft is placed in PTC to
stabilize the thermal response by spacecraft structures and systems.

At 30:40:49 g.e.t., a midcourse correction maneuver was made using the
service module propulsion system.  The crew preparations for the burn and
the burn itself were monitored by the Mission Control Center (MCC) at MSC
by telemetered data and by television from the spacecraft. This midcourse
correction maneuver was a 23.2 feet per second hybrid 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 Moon to insure the proper lighting conditions at the
landing site. Apollo 8, 10, and 11 flew a pure free-return trajectory
until lunar orbit insertion. The Apollo 13 hybrid transfer maneuver
lowered the predicted closest approach, or pericynthion, altitude at the
Moon from 210 to 64 nautical miles. 

     From launch through the first 46 hours of the mission, the perform-
ance 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 caused by either a short circuit or an open
circuit in the gage wiring or a short circuit between the gage plates.
Subsequent events indicated that a short was the more likely failure mode. 

     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 LM
to commence inflight inspection of the LM. Ground tests 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 LM for pressurizing propellant tanks. 

     The LM was powered down and preparations were underway to close the
LM hatch and run through the presleep checklist when the accident in
oxygen tank no. 2 occurred. 

     At 55:52:30 g.e.t., a master alarm on the CM caution and warning
system alerted the crew to a low pressure indication in the cryogenic
hydrogen tank no. 1. This tank had reached the low end of its normal
operating pressure range several times previously during the flight.
At 55:52:58, flight controllers in the MCC requested the crew to turn
on the cryogenic system fans and heaters.

     The Command Module Pilot (CMP) acknowledged the fan cycle request at
55:53:06 g.e.t., and data indicate that current was applied to the oxygen
tank no. 2 fan motors at 55:53:20. 

     About 1-1/2 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. 1 or its lines or valves.
Photos of the SM taken by the crew later in the mission show the panel
missing, the fuel cells on the shelf above the oxygen shelf tilted, and
the high-gain antenna damaged. 

     The resultant loss of oxygen made the fuel cells inoperative, leav-
ing 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 vent-
ing). 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. 


TABLE 4-II.- DETAILED CHRONOLOGY FROM
2.5 MINUTES BEFORE THE ACCIDENT TO 5 MINUTES AFTER THE 
ACCIDENT

Time, g.e.t.                       Event

      Events During 52 Seconds Prior to First Observed Abnormality

55:52-:31	Master caution and warning triggered by low hydrogen
              	pressure in tank no. 1. Alarm is turned off after
              	4 seconds.
     
55:52:58        Ground requests tank stir.
     
55:53:06        Crew acknowledges tank stir.

55:53:18        Oxygen tank no. 1 fans on.
     
55:53:19        Oxygen tank no. 1 pressure decreases 8 psi.
     
55:53:20        Oxygen tank no. 2 fans turned on.
     
55:53:20        Stabilization control system electrical disturbance
         	indicates a power transient.
     
55:53:21     	Oxygen tank no. 2 pressure decreases 4 psi.

      Abnormal Events During 90 Seconds Preceding the Accident

55:53:22.718	Stabilization control system electrical distrubance 
		indicates a power transient.

55:53:22.757  	1.2-volt decrease in ac bus 2 voltage.
     
55:53:22.772    11.1-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    11-volt decrease in ac bus 2 voltage for one
               	sample.
     
55:53:38.085    Stabilization control system electrical disturbance
               	indicates a power transient.
     
55:53:41.172    22.9-amp rise in fuel cell 3 current for one sample.
     
55:53:41.192    Stabilization control system electrical disturbance
              	indicates a power transient.
     
55:54:00       	Oxygen tank no. 2 pressure rise ends at a pressure
               	of 953.8 psia.
     
55:54:15        Oxygen tank no. 2 pressure begins to rise.
     
55:54:30       	Oxygen tank no. 2 quantity drops from full scale
               	for 2 seconds and then reads 75.3 percent.
    
55:54:31        Oxygen tank no. 2 temperature begins to rise
               	rapidly.
     
55:54:43      	Flow rate of oxygen to all three fuel cells begins
               	to decrease.
     
55:54:45      	Oxygen tank no. 2 pressure reaches maximum value
               	of 1008.3 psia.
     
55:54:48      	Oxygen tank no. 2 temperature rises 40 F for one
               	sample (invalid reading).
     
55:54:51        Oxygen tank no. 2 quantity Jumps to off-scale high
               	and then begins to drop until the time of telemetry
               	loss, indicating failed sensor.
     
55:54:52        Oxygen tank no. 2 temperature reads -151.3 F.
     
55:54:52.703    Oxygen tank no. 2 temperature suddenly goes off
               	scale low, indicating failed sensor.
     
55:54:52.763    Last telemetered pressure from oxygen tank no. 2
               	before telemetry loss is 995.7 psia.
     
55:54:53.182    Sudden accelerometer activity on X, Y, and Z axes.
     
55:54:53.220    Stabilization control system body rate changes
               	begin.
     
55:54:53.323    Oxygen tank no. 1 pressure drops 4.2 psi.
     
55:54:53.5     	2.8-amp rise in total fuel cell current.
     
55:54:53.542    X, Y, and Z accelerations in CM indicate 1.17g,
               	0.65g and 0.65g, respectively.

             1.8-Second Data Loss
     
55:54:53.555	Loss of telemetry begins.
     
55:54:53.555+	Master caution and warning triggered by dc main
               	bus B undervoltage. 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.
     
55:54:54.741    Nitrogen pressure in fuel cell 1 is off-scale low
               	indicating failed sensor.
     
55:54:55.35     Recovery of telemetry data.

           Events During 5 Minutes Following the Accident

55:54:56	Service propulsion system engine valve body temperature 
		begins a rise of 1.65 F in 7 seconds. 
     
55:54:56     	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.
     
55:54:56     	Total fuel cell current is 15 amps higher than the
               	final value before telemetry loss. High current
               	continues for 19 seconds.
     
55:54:56    	Oxygen tank no. 2 temperature reads off-scale high
               	after telemetry recovery, probably indicating failed
               	sensors.
     
55:54:56      	Oxygen tank no. 2 pressure reads off-scale low fol-
               	lowing telemetry recovery, indicating a broken supply
               	line, a tank pressure below 19 psi, or a failed sensor.
     
55:54:56      	Oxygen tank no. 1 pressure reads 781.9 psia and
               	begins to drop steadily.
     
55:54:57   	Oxygen tank no. 2 quantity reads off-scale high
               	following telemetry recovery indicating failed sensor.
     
55:54:59     	The reaction control system helium tank C temperature
               	begins a 1.66 F increase in 36 seconds.
     
55:55:01        Oxygen flow rates to fuel cells 1 and 3 approached
               	zero after decreasing for 7 seconds.
     
55:55:02       	The surface temperature of the service module oxi-
               	dizer tank in bay 3 begins a 3.8 F increase in a
               	15-second period.
     
55:55:02      	The service propulsion system helium tank temperature
               	begins a 3.8 F increase in a 32-second period.
    
55:55:09      	Dc main bus A voltage recovers to 29.0 volts, dc
               	main bus B recovers to 28.8 volts.
     
55:55:20     	Crew reports, "I believe we've had a problem here."
     
55:55:35     	Crew reports, "We've had a main B bus undervolt."
     
55:55:49       	Oxygen tank no. 2 temperature begins steady drop
               	lasting 59 seconds, probably indicating failed sensor.
     
55:56:10      	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."
     
55:56:38     	Oxygen tank no. 2 quantity becomes erratic for 69
               	seconds before assuming an off-scale-low state,
               	indicating failed sensor.
     
55:57:04       	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."
     
55:57:39     	Master caution and warning triggered by dc main
               	bus B undervoltage. Alarm is turned off in
               	6 seconds.
     
55:57:40      	Dc main bus B drops below 26.25 volts and continues
               	to fall rapidly.
     
55:57:44        Ac bus 2 fails within 2 seconds
     
55:57:45        Fuel cell 3 fails.
     
55:57:59       	Fuel cell 1 current begins to decrease.
     
55:58:02      	Master caution and warning caused by ac bus 2
               	being reset. Alarm is turned off after 2 seconds.
     

55:58:06      	Master caution and warning triggered by dc main
               	bus A undervoltage. Alarm is turned off in 13
               	seconds.
     

55:58:07     	Dc main bus A drops below 26.25 volts and in the
               	next few seconds levels off at 25.5 volts .
     
55:58:07       	Crew reports, "ac 2 is showing zip."

55:58:25      	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."
     
56:00:06       	Master caution and warning triggered by high hydrogen
               	flow rate to fuel cell 2. Alarm is turned off in
               	2 seconds.


PART 4.  SUMMARY ANALYSIS OF THE ACCIDENT


     Combustion in oxygen tank no. 2 led to failure of that tank, damage
to oxygen tank no. 1 or its lines or valves adjacent to tank no. 2,
removal of the bay 4 panel and, through the resultant loss of all three
fuel cells, to the decision to abort the Apollo 13 mission. In the attempt
to determine the cause of ignition in oxygen tank no. 2, the course of
propagation of the combustion, the mode of tank failure, and the way in
which subsequent damage occurred, the Board has carefully sifted through
all available evidence and examined the results of special tests and
analyses conducted by the Apollo organization and by or for the Board
after the accident. (For more information on details of mission events,
design, manufacture and test of the system, and special tests and analyses
conducted in this investigation, refer to Appendices B, C, D, E, and F of
this report.)

     Although tests and analyses are continuing, sufficient information is
now available to provide a reasonably clear picture of the nature of the
accident and the events which led up to it. It is now apparent that the
extended heater operation at KSC damaged the insulation on wiring in the
tank and thus made the wiring susceptible to the electrical short circuit
which probably initiated combustion within the tank. While the exact point
of initiation of combustion may never be known with certainty, the
nature of the occurrence is sufficiently understood to permit taking
corrective steps to prevent its recurrence. 

The Board has identified the most probable failure mode.

     The following discussion treats the accident in its key phases:
initiation, propagation of combustion, loss of oxygen tank no. 2 system
integrity, and loss of oxygen tank no. 1 system integrity. 

INITIATION

Key Data

55:53:20*	Oxygen tank no. 2 fans turned on.

55:53:22.757 	1.2-volt decrease in ac bus 2 voltage.

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


55:53:22.772   	11.1-ampere "spike" recorded in fuel cell 3 current
               	followed by drop in current and rise in voltage typ-
               	ical of removal of power from one fan motor--indicat-
               	ing 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 11.1-ampere current spike and
simultaneously a voltage-drop spike were recorded in the spacecraft
electrical system. Immediately thereafter, current drawn from the fuel
cells decreased by an amount consistent with the loss of power to one fan.
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 low-level combustion at
the time. These facts point to the likelihood that an electrical short
circuit with arcing occurred in the fan motor or its leads to initiate the
accident sequence. The energy available from the short circuit was
probably 10 to 20 joules. Tests conducted during this investigation have
shown that this energy is more than adequate to ignite Teflon of the
type contained within the tank. (The quantity gage in oxygen tank no. 2
had failed at 46:40 g.e.t. 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 dur-
ing the abnormal detanking operation at KSC prior to launch. 

     Furthermore, there is no evidence pointing to any other mechanism of
initiation. 

PROPAGATION OF COMBUSTION

Key Data

55:53:36    	Oxygen tank no. 2 pressure begins rise (same event
               	noted previously).

55:53:38.057    11-volt decrease recorded in ac bus 2 voltage.

55:53:41.172    22.9-ampere "spike" recorded in fuel cell 3 current,
            	followed by drop in current and rise in voltage typ-
	        ical 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 behav
               	ior indicates the gage short circuit may have cor-
               	rected 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 tem-
perature, 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 supercrit-
ical oxygen in zero-g are extremely limited, those which are available
indicate that the rate of combustion is generally consistent with these
observations. The cause of the 15-second period of relatively constant
pressure first indicated at 55:53:59.763 has not been precisely deter-
mined; it is believed to be associated with a change in reaction rate as
combustion proceeded through various Teflon elements. 

     While there is enough electrical power in the tank to cause ignition
in the event of a short circuit or abnormal heating in defective wire,
there is not sufficient electric power to account for s11 of the energy
required to produce the observed pressure rise.

LOSS OF OXYGEN TANK NO. 2 SYSTEM INTEGRITY

				Key Data

55:54:52      	Last valid temperature indication (-151 F) from
               	oxygen tank no. 2.
55:54:52.763    Last pressure reading from oxygen tank no. 2 before
               	loss of data--996 psia.
55:54:53.182   	Sudden accelerometer activity on X, Y, and Z axes.
55:54:53.220   	Stabilization control system body rate changes begin.
55:54:53.555*   Loss of telemetry data begins.
55:54:55.35     Recovery of telemetry data.
55:54:56        Various temperature indications in SM begin slight
               	rises.
55:54:56      	Oxygen tank no. 2 temperature reads off-scale high.
55:54:56      	Oxygen tank no. 2 Pressure reads off-scale low.

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

     After the relatively slow propagation process described above took
place, there was a relatively abrupt loss of oxygen tank no. 2 integ-
rity. 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 all
signals from the spacecraft were lost about 1.85 seconds after the last
presumably valid reading from within the tank, a temperature reading,
and 0.8 second after the last presumably valid pressure reading (which may
or may not reflect the pressure within the tank itself since the pressure
transducer is about 20 feet of tubing length distant). Abnormal spacecraft
accelerations were recorded approximately 0.42 second after the last
pressure reading and approximately 0.38 second before the loss of signal.
These facts all point to a relatively sudden loss of integrity. At about
this time, several solenoid valves, including the oxygen valves feeding
two of the three fuel cells, were shocked to the closed position. The
"bang" reported by the crew also probably occurred in this time period.
Telemetry signals from Apollo 13 were lost for a period of 1.8 seconds.
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 infor-
mation, analyses, and tests performed during this investigation indicate
that most probably the combustion within the pressure vessel ultimately
led to localized heating and failure at the pressure vessel closure. It is
at this point, the upper end of the quantity probe, that the l/2-inch
Inconel conduit is located, through which the Teflon-insulated wires enter
the pressure vessel. It is likely that the combustion progressed along the
wire insulation and reached this location where all of the wires come
together. This, possibly augmented by ignition of the metal in the upper
end of the probe, led to weakening and failure of the closure or the
conduit, or both. 

     Failure at this point would lead immediately to pressurization of the
tank dome, which is equipped with a rupture disc rated at about 75 psi.
Rupture of this disc or of the entire dome would then release oxygen,
accompanied by combustion products, into bay 4. The accelerations
recorded were probably caused by this release. 

     Release of the oxygen then began to pressurize the oxygen shelf space
of bay 4. 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 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 0F OXYGEN TANK NO. 1 INTEGRITY

                                Key Data

55:54:53.323 		Oxygen tank no. 1 pressure drops 4 psia (from 883 
			psia to 879 psia).
55:54:53.555 to     	Loss of telemetry data.
55:54:55.35     
55:54:56 	        Oxygen tank no. 1 pressure reads 782 psia and drops
          		steadily. Pressure drops over a period of 130 min- 
          		utes 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. 1, it is presumed that the loss of oxygen
tank no. 1 integrity resulted from the oxygen tank no. 2 system failure.
The relatively sudden, and possibly violent, event associated with loss of
integrity of the oxygen tank no. 2 system could have ruptured a line to
oxygen tank no. 1, or have caused a valve to leak because of mechanical
shock. 


PART 5. APOLLO-13 RECOVERY

UNDERSTANDING  THE PROBLEM

     In the period immediately following the caution and warning alarm for
main bus B undervoltage, and the associated "bang" reported by the crew,
the cause of the difficulty and the degree of its seriousness were not
apparent. 

     The 1.8-second loss of telemetered data was accompanied by the
switching of the CSM high-gain antenna mounted on the SM adjacent to bay 4
from narrow beam width to wide beam width. 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. 

The Lunar Module Pilot was in the tunnel between the CSM and the LM,
returning to the CSM. The command Module Pilot was in the left-hand
couch, monitoring spacecraft performance. Because of the master alarm
indicating low voltage, the CMP moved across to the right-hand couch where
CSM voltages can be observed. He reported that voltages were "looking
good" at 55:56:10. At this time, main bus B had recovered and fuel cell 3
did not fail for another 1-1/2 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 CM panel arrangement). 

     When fuel cells 1 and 3 electrical output readings went to zero, the
ground controllers could not be certain that the cells had not somehow
been disconnected from their respective 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 config-
uration was known. When it was realized that fuel cells 1 and 3 were not
functioning, the crew was directed to perform an emergency powerdown to
lower the load on the remaining fuel cell. Observing the rapid decay in
oxygen tank no. 1 pressure, controllers asked the crew to switch power to
the oxygen tank no. 2 instrumentation. When this was done, and it was
realized that oxygen tank no. 2 had failed, the extreme seriousness of the
situation became clear. 

     During the succeeding period, efforts were made to save the remain-
ing oxygen in the oxygen tank no. 1. Several attempts were made, but had
no effect. The pressure continued to decrease. 

     It was obvious by about 1-1/2 hours after the accident that the
oxygen tank no. 1 leak could not be stopped and that shortly it would be
necessary to use the LM as a lifeboat' for the remainder of the mission. 

     By 58:40 g.e.t., the LM had been activated, the inertial guidance
reference transferred from the CSM guidance system to the LM guidance
system, and the CSM-systems were turned off. 

     RETURN TO EARTH

     The remainder of the mission was characterized by two main activ-
ities--planning and conducting the necessary propulsion maneuvers to
return the spacecraft to Earth, and managing the use of consumables in
such a way that the LM, which is designed for a basic mission with two
crewmen for a relatively short duration, could support three men and serve
as the actual control vehicle for the time required. 

     One significant anomaly was noted during the remainder of the 
mission.  At about 97 hours 14 minutes into the mission, the LMP reported 
hearing a "thump" and observing venting from the LM.  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 afater 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 hours 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 (LiOH) filter 
cannisters used to remove carbon dioxide from the spacecraft 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 LiOH cannisters in the LM  cabin atmosphere cleaning system 
(see fig. 4-16).  At splashdown, many hours of each consumable remained 
available (see figs. 4-17 through 4-19 and table 4-III). 


TABLE 4-III.--CABIN ATMOSPHERE CARBON DIOXIDE REMOVAL BY LITHIUM HYDROXIDE

Required                      85 hours

Available in LM               53 hours

Available in CM              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.


CHAPTER 5
FINDINGS, DETERMINATIONS, AND RECOMMENDATIONS

PART 1. 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 con-
tractors, and on an extensive series of special tests and analyses per-
formed 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.  Ac-
cordingly, the Board has concentrated on this tank; on its design, manu-
facture, 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 over-
all direction of MSC, will continue to generate new information relative
to this accident. It is possible that this evidence may lead to conclu-
sions 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 organi-
zational elements responsible for it. It was found that the accident was
not the result of a chance malfunction in a statistical sense, but rather
resulted from an unusual combination of mistakes, coupled with a somewhat
deficient and unforgiving design. In brief, this is what happened: 

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

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

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

     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 KSC subjected the tank to an
extended period of heater operation and pressure cycling. These proce-
dures 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, ER, and Beech did not recognize the possibility of damage due to
overheating. Many of these officials were not aware of the extended heater
operation. In any event, adequate-thermostatic switches might have been
expected to protect the tank. 

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

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

     j. The thermostatic switches could accommodate the 65 V dc during
tank pressurization because they normally remained cool and closed.  How-
ever, they could not open without damage with 65 V dc power applied. They
were never required to do so until the special detanking. During this
procedure, as the switches started to open when they reached their upper
temperature limit, they were welded permanently closed by the resulting
arc and were rendered inoperative as protective thermostats. 

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

     l. As shown by subsequent tests, failure of the thermostatic switches
probably permitted the temperature of the heater tube assembly to reach
about 1000 F in spots during the continuous 8-hour period of heater
operation. Such heating has been shown by tests to severely damage the
Teflon insulation on the fan motor wires in the vicinity of the heater
assembly. From that time on, including pad occupancy, the oxygen tank no.
2 was in a hazardous condition when filled with oxygen and electrically
powered. 

     m. It was not until nearly 56 hours into the mission, however, that
the fan motor wiring, possibly moved by the fan stirring, short circuited
and ignited its insulation by means of an electric arc. The resulting
combustion in the oxygen tank probably overheated and failed the wiring
conduit where it enters the tank, and possibly a portion of the tank it-
self. 

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

     The accident is judged to have been nearly catastrophic. Only out-
standing performance on the part of the crew, Mission Control, and other
members of the team which supported the operations successfully returned
the crew to Earth

     In investigating the accident to Apollo 13, the Board has also
attempted to identify those additional technical and management lessons
which can be applied to help assure the success of future space flight
missions: several recommendations of this nature are included. 

     The Board recognizes that the contents of its report are largely of
a critical nature. The report highlights in detail faults or deficiencies
in equipment and procedures that the Board has identified. This is the
nature of a review board report.


     It is important, however, to view the criticisms in this report in
a broader context. The Apollo spacecraft system is not without short-
comings, but it is the only system of its type ever built and success-
fully 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 accomplish-
ments speak for themselves. By hardheaded self-criticism and continued
dedication, this team can maintain this nation's preeminence in space.

PART 2. ASSESSMENT OF ACCIDENT

FAILURE OF OXYGEN TANK NO. 2

1. 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. 1, and
a resulting loss of power from the oxygen-fed fuel cells.

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

c. Oxygen tank no. 2 contained materials, including Teflon and
aluminum, which if ignited 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 thermo-
static switches on the heaters were required to open while
powered by 65 V dc in order to protect the heaters from over-
heating. 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. 1 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 immediately prior to the accident indicate
electrical disturbances of a character which would be caused
by short circuits accompanied by electrical arcs in the fan
motor or its leads in oxygen tank no. 2.

i. The pressure and temperature within oxygen tank no. 2 rose
abnormally during the 1-1/2 minutes immediately prior to the
accident.

Determinations

(1) 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.

(5) 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 tem-
peratures 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 sub-
sequent tests. These tests also verified that the l-ampere
fuses on the fan motors would pass sufficient energy to ig-
nite 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 over-
heat and fail the tank or its associated tubing. The possi-
biIity of such failure at the top of the tank was demon-
strated by subsequent tests.

(8) The rate of flame propagation along Teflon-insulated wires
as measured in subsequent tests is consistent with the in-
dicated rates of pressure rise within the tank.

SECONDARY EFFECTS OF TANK FAILURE

2. Findings


a. Failure-of the tank was accompanied by several events in-
cluding:

A 'bang" as heard by the crew.

Spacecraft motion as felt by the crew and as measured by
the attitude control system and the accelerometers in the
command module (CM).

Momentary loss of telemetry.

Closing of several valves by shock loading.

Loss of integrity of the oxygen tank no. 1 system.

Slight temperature increases in bay 4 and adjacent sectors
of the SM.

Loss of the panel covering bay 4 of the SM, as observed and
photographed by the crew.

Displacement of the fuel cells as photographed by the crew.

Damage to the high-gain antenna as photographed by the crew.

b. The panel covering of bay 4 could be blown off by pressuri-
zation of the bay. About 25 psi of uniform pressure in bay 4
is required to blow off the panel.

c. 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 10 psi on the CM heat shield and this would separate
the CM from the SM.

Determinations

(1) Failure of the oxygen tank no. 2 caused a rapid local
pressurization of bay 4 of the SM by the high-pressure
oxygen that escaped from the tank.  This pressure pulse may
have blown off the panel covering bay 4. 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 sub-
sequent tests.

(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. 1 and the subsequent
loss of power resulted from the tank no. 2 failure.

(7) Telemetry, although good, is insufficient to pin down the
exact nature, sequence, and location of each event of the
accident in detail.

(8) The telemetry data, crew testimony, photographs, and special
tests and analyses already completed are sufficient to under-
stand 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

(1) 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
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.

(4) The amount of material in the tank which could be ignited
and burned in the given environment could have been reduced
significantly.

t5) 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 atten-
tion to the danger of fire in a pure oxygen environment,
failed to recognize the deficiencies of the tank.

     PREFLIGHT DAMAGE TO TANK WIRING

4. Findings

a. The oxygen tank no. 2 heater assembly contained two thermo-
static switches designed to protect the heaters from over-
heating.

b. The thermostatic switches were designed to open and interrupt
the heater current at 80 + 10 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 Cor-
poration 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 inter-
rupt this load. 

f. 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
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 be-
cause this would only happen if the oxygen supply in a tank
were depleted to nearly zero.

h. 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, speci-
fied allowable heater voltages and currents were not exceeded.

1. 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

(1) 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.

(3) The fact that the switches were not rated to open at 65 V dc
was not detected by NASA, ER, or Beech in their reviews of
documentation or in qualification and acceptance testing.

(4) The failed switches resulted in severe overheating. Subse-
quent tests showed that heater assembly temperatures could
have reached about 1000 F.

(5) The high temperatures severely damaged the Teflon insulation
on the wiring in the vicinity of the heater assembly and set
the stage for subsequent short circuiting. As shown in
subsequent tests, this damage could range from cracking to
total oxidation and disappearance of the insulation.

(6) During and following the special detanking, the oxygen tank
no. 2 was in a hazardous condition whenever it contained
oxygen and was electrically energized.

PART 3. SUPPORTING CONSIDERATIONS

DESIGN, MANUFACTURING, AND TEST

5. Finding

The pressure vessel of the supercritical oxygen tank is con-
structed of Inconel 718, and is moderately stressed at normal
operating pressure.

Determination

From a structural viewpoint, the supercritical oxygen pressure
vessel is quite adequately designed, employing a tough material
well chosen for this application. The stress analysis and the
results of the qualification burst test program confirm the
ability of the tank to exhibit adequate performance in its in-
tended application.

6. Findings

a. The oxygen tank design includes two unsealed electric fan
motors immersed in supercritical oxygen.

b. Fan motors of this design have a test history of failure
during acceptance test which includes phase-to-phase and
phase-to-ground faults.

c. The fan motor stator windings are constructed with Teflon-
coated, ceramic-insulated, number 36 AWG wire. Full phase-
to-phase and phase-to-ground insulation is not used in the
motor design.

d. The motor case is largely aluminum.

Determinations

(1) 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 ques-
tionable practice.

7 Findings

a. The cryogenic oxygen storage tanks contained materials that could be
ignited and which will burn under the conditions prevailing within the
tank, including Teflon, aluminum, solder, and Drilube 822. 

b. The tank contained electrical wiring exposed to the super-
critical oxygen. The wiring was insulated with Teflon.

c. Some wiring was in close proximity to heater elements and
to the rotating fan.

d. 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 cryo-
genic oxygen storage tanks of the SM was exposed to rela-
tively sharp metal edges of tank inner parts during manu-
facturing assembly operations.

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

Determinations

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

(2) Scraping of the electrical wiring insulation against metal
inner parts of the tank constituted a substantial cumulative
hazard during assembly, handling, test, checkout, and opera-
tional use.

(3) "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 po-
tential 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 man-
ufacturing.

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 com-
pensate 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.

10. 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 combina-
tion of high-pressure oxygen, Teflon, and electrical wiring
as in the oxygen tank no. 2. The wiring is unfused and is
routed through a 10-amp circuit breaker.

Determination

The fuel cell oxygen supply valve module has been identified as
potentially hazardous.

11. Findings

a. In the normal sequence of cryogenic oxygen storage tank in-
tegration 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.

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 com-
pensate 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.

10. 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 combina-
tion of high-pressure oxygen, Teflon, and electrical wiring
as in the oxygen tank no. 2. The wiring is unfused and is
routed through a 10-amp circuit breaker.

Determination

The fuel cell oxygen supply valve module has been identified as
potentially hazardous. 

11. Findings

a. In the normal sequence of cryogenic oxygen storage tank in-
tegration 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.

b. The SM undergoes factory transportation, air shipment to KSC,
and subsequent ground transportation and handling.

Determination

There were environments during the normal sequence of operations
subsequent to the final acceptance tests at Beech that could
cause a loose-fitting set of fill tube parts to become displaced.

12. Findings

a. At North American Rockwell, Downey, California, in the
attempt to remove the oxygen shelf assembly from SM 106,
a bolt restraining the inner edge of the shelf was not re-
moved.

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 XTA0008 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 empha-
sizing electrical continuity of internal wiring were accom-
plished before reinstallation.

Determinations

(1) Displacement of fill tube parts could have occurred, during
the "shelf drop" incident at the prime contractor's plant,
without detection.

(2) Other damage to the tank may have occurred from the jolt,
but special tests and analyses indicate that this is un-
likely.

(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.

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 pro-
tracted 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.

f. Tests subsequent to the flight showed that the current
associated with the KSC 65 V dc ground powering of the
heaters would cause the thermostatic switch contacts to
weld closed if they attempted to interrupt this current.

g. A second test showed that without functioning thermostatic
switches, temperatures in the 800 to 1000 F range would
exist at locations on the heater tube assembly that were in
close proximity with the motor wires. These temperatures
are high enough to damage Teflon and melt solder.

Determinations

(1) Oxygen tank no. 2 (XTA 0008) did not detank after CDDT in a
manner comparable to its performance the last time it had
contained liquid oxygen, i.e., in acceptance test at Beech.

(2) Such evidence indicates that the tank had undergone some
change of internal configuration during the intervening
events of the previous 3 years.

(3) The tank conditions during the special detanking procedures
were outside all prior testing of Apollo CSM cryogenic oxygen
storage tanks. Heater assembly temperatures measured in sub-
sequent tests exceeded 1000 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 un-
supported 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 re-
garding the cryogenic oxygen system included incomplete and
inaccurate information.

d. The MSC Test Specification Criteria Document (TSCD) which
was used by KSC in preparing detailed tank test procedures
states the tank allowable heater voltage and current as 65
to 85 V dc and 9 to 17 amperes with no restrictions on time.

Determinations

(1) NR and MSC personnel who prepared 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.

(3) Launch operations personnel at KSC stayed within the
specified tank heater voltage and current limits during the
detanking at KSC.

16. Findings

a. After receipt of the Block 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 compati-
bility between the switch and the GSE.

d. No tests were specified by MSC, NR, or Beech to check this
switch under load.

Determinations

(1) NR and Beech specifications governing the powering and the
thermostatic switch protection of the heater assemblies were
inadequate.

(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

(1) MSC, KSC, and NR personnel did not know that the tank heater
thermostatic switches did not protect the tank from over-
heating.

(2) If the long period of continuous heater operation with failed
thermostatic switches had been known, the tank would have
been replaced.

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.

I9. 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 THROUGH ACCIDENT

20. Findings

a. The center engine of the S-II stage of the Saturn V launch
vehicle prematurely shut down at 132 seconds due to large
16 hertz oscillations in thrust chamber pressure.

b. Data indicated less than 0.lg vibration in the CM.

Determinations

(1) Investigation of this S-II anomaly was not within the purview
of the Board except insofar as it relates to the Apollo 13
accident.

(2) The resulting oscillations or vibration of the space vehicle
probably did not affect the oxygen tank.

21. Findings

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

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

Determinations

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

(2) The short circuit could have been caused by either a com-
pletely 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 1 ampere at 55:53:19.

c. The oxygen tank no. 1 pressure decreased 8 psi at 55:53:19
due to normal destratification.

Determination

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

23. Findings

a. Spacecraft current increased by L-1/2 amperes and ac bus 2
voltage decreased 0.6 volt at 55:53:20.

b. Stabilization and Control System (SCS) gimbal command telem-
etry channels, which are sensitive indicators of electrical
transients associated with switching on or off of certain
spacecraft electrical loads, showed a negative initial tran-
sient during oxygen tank no. 2 fan turnon cycles and a posi-
tive initial transient during oxygen tank no. 2 fan turnoff
cycles during the Apollo 13 mission. A negative initial
transient was measured in the SCS at 55:53:20.

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

Determinations

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

(2) It cannot be determined whether or not they were rotating
because the pressure decrease was too small to conclusively
show destratification. It is likely that they were.

24. Finding

An 11.1-amp spike in fuel cell 3 current and a momentary
1.2-volt decrease were measured in ac bus 2 at 55:53:23.

Determinations

(1) A short circuit occurred in the circuits of the fans in
oxygen tank no. 2 which resulted in either blown fuses or
opened wiring, and one fan ceased to function.

(2) The short circuit probably dissipated an energy in excess
of 10 joules which, as shown in subsequent tests, is more
than sufficient to ignite Teflon wire insulation by means
of an electric arc.

25. Findings

a. A momentary 11-volt decrease in ac bus 2 voltage was
measured at 55:53:38.

b. A 22.9-amp spike in fuel cell 3 current was measured at
55:53:41.

c. After the electrical transients, CM current and ac bus 2
voltage returned to the values indicated prior to the turn-
on of the fans in oxygen tank no. 2.

Determination

Two short circuits occurred in the oxygen tank no. 2 fan cir-
cuits between 55:53:38 and 55:53:41 which resulted in either
blown fuses or opened wiring, and the second fan ceased to
function.

26. Finding

Oxygen tank no. 2 telemetry showed a pressure rise from 887 to
954 psia between 55:53:36 and 55:54:00. 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

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

(2) Since no other known energy source in the tank could produce
this pressure buildup, it is concluded to have resulted from
combustion initiated by the first short circuit which started
a wire insulation fire in the tank.

27. Findings

a. The pressure relief valve was designed to be fully open at
about 1000 psi.

b. Oxygen tank no. 2 telemetry showed a pressure drop from
1008 psia at 55:54:45 to 996 psia at 55:54:53, at which time
telemetry data were lost.

Determination

This drop resulted from the normal operation of the pressure
relief valve as verified in subsequent tests.

28. Findings

a. At 55:54:29, when the pressure in oxygen tank no. 2 exceeded
the master caution and warning trip level of 975 psia, the CM
master alarm was inhibited by the fact that a warning of low
hydrogen pressure was already in effect, and neither the crew
nor Mission Control was alerted to the pressure rise.

b. The master caution and warning system logic for the cryogenic
system is such that an out-of-tolerance condition of one
measurement which triggers a master alarm prevents another
master alarm from being generated when any other parameter in
the same system becomes out-of-tolerance.

c. The low-pressure trip level of the master caution and warning
system for the cryogenic storage system is only 1 psi below
the specified lower limit of the pressure switch which con-
trols the tank heaters. A small imbalance in hydrogen tank
pressures or a shift in transducer or switch calibration can cause 
the master caution and warning to be triggered pre-
ceding 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 one Mission Control did not observe it.

Determinations 

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

(2) With the present master caution and warning system, a space-
craft problem can go unnoticed because of the presence of a
previous out-of-tolerance condition in the same subsystem.

(3) Although a master alarm at 55:54:29 or observance of a limit
sense light in Mission Control could have alerted the crew
or Mission Control in sufficient time to detect the pressure
rise in oxygen tank no. 2, no action could have been taken
at that time to prevent the tank failure. However, the in-
formation 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

(1) 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.

30. Finding

Oxygen tank no. 2 telemetry indicated the following changes:
(1) quantity decreased from off-scale high to off-scale low in
2 seconds at 55:54:30, (2) quantity increased to 75.3 percent at
55:54:32, and (3) quantity was off-scale high at 55:54:51 and
later became erratic.

Determinations

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

(2) Immediately preceding and following this time period, the
indications were caused by electrical faults.

31. Findings

a. At about 55:54:53, or about half a second before telemetry
loss, the body-mounted linear accelerometers in the command
module, which are sampled at 100 times per second, began
indicating spacecraft motions. 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 milli-
seconds before data loss.

b. The body-mounted roll, pitch, and yaw rate gyros showed low-
level 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 space-
craft between 55:54:53 and 55:54:55.

d. Doppler tracking data measured an incremental velocity com-
ponent of 0.26 fps along a line from the Earth to the space-
craft, 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.

Determinations

(1) The spacecraft was subjected to abnormal forces at approxi-
mately 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 place in the SM at approximately the time of panel
separation.

33. Findings

a. The telemetered nitrogen pressure in fuel cell 1 was off-
scale low at reacquisition of data at 55:54:55.

b. Fuel cell 1 continued to operate for about 3 minutes past
this time.

c. The wiring to the nitrogen sensor passes along the top of
the shelf which supports the fuel cells immediately above
the oxygen tanks.

Determinations

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

(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. 1 pressure decreased rapidly from 879 psia to
782 psia at approximately 55:54:54 and then began to decrease
more slowly at 55:54:56.


Determination

A leak caused loss of oxygen from tank no. 1 beginning at approxi-
mately 55:54:54.

35. Findings

a. Oxygen flow rates to fuel cells 1 and 3 decreased in a
5-second period beginning at 55:54:55, but sufficient volume
existed in lines feeding the fuel cells to allow them to
operate about 3 minutes after the oxygen supply valves were
cut off.

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 magnitude as
the shock required to close the RCS valves. 

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

Determination

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

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 indica-
tions were as follows

Oxygen tank no. 1: Pressure: Several hundred psi below
normal

Quantity-Normal

Temperature: Normal


Oxygen tank no. 2: Pressure: Off-scale low

Quantity: Off-scale high

Temperature: Off-scale high

c. The nitrogen pressure in fuel cell 1 indicated zero, which was
incompatible with the 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 nitro-
gen 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 in-
vestigation.

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 credita-
bility of the telemetered data.

37. Findings

a. During the 3 minutes following data loss, neither the flight
controllers nor the crew noticed the oxygen flows to fuel
cells 1 and 3 were less than 0.1 lb./hr. These were unusually
low readings for the current being drawn.

b. Fuel cells 1 and 3 failed at about 3 minutes after the data
loss.

c. After the fuel cell failures, which resulted in dc main
bus B failure and the undervoltage condition on dc main bus A,
Mission Control diverted its prime concern from what was
initially believed to be a cryogenic system instrumentation
problem to the electrical power system.

d. Near-zero oxygen flow to fuel cells 1 and 3 was noted after
the main bus B failure, but this was consistent with no power
output from the fuel cells.

e. The flight controllers believed that the fuel cells could
have been disconnected from the busses and directed the crew
to connect fuel cell 1 to dc main bus A and fuel cell 3 to
dc main bus B.

f. The crew reported the fuel cells were configured as directed
and that the talkback indicators confirmed this.

Determinations

(1) Under these conditions it was logical for the flight con-
trollers 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 space-
craft do not indicate closed unless both the hydrogen and oxygen
valves are closed.

Determinations

(1) 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 1 and 3, although they would never-
theless have failed when oxygen tank no. 1 ceased to supply
oxygen.


(2) The ultimate outcome would not have been-changed, but had the
fuel cells not failed, Mission Control and the crew would not
have had to contend with the failure of dc main bus B and ac
bus 2 or attitude control problems while trying to evaluate
the situation.

Reaction Control System

39. Findings

a. 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 1-1/2-hour period following the accident, Mission
Control noted that SM RCS quad C propellant was not being
used, although numerous firing signals were being sent to it.

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 1-1/2-hour period immediately following the
accident Mission Control advised the crew which SM RCS
thrusters to Power and which ones to unpower.

Determinations

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

Helium isolation valves in quads B and D
Secondary fuel pressurization valves in quads A and C

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

(3) Mission Control correctly determined the status of the RCS
system and properly advised the crew on how to regain auto-
matic attitude control.

Management of Electrical System

40. Findings

a. After fuel cell l failed, the total dc main bus A load was
placed on fuel cell 2 and the voltage dropped to approxi-
mately 25 volts, causing a caution and warning indication
and a master alarm.

b. After determining the fuel cell 2 could not supply enough
power to dc main bus A to maintain adequate voltage, the crew
connected entry battery A to this bus as an emergency measure
to increase the bus voltage to its normal operating value.

c. Mission Control directed the crew to reduce the electrical
load on dc main bus A by following the emergency powerdown
checklist contained in the onboard Flight Data File.

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

e. Mission Control then directed the crew to charge this battery
in order to get as much energy back into it as possible,
before the inevitable loss of the one functioning fuel cell.

Determinations

(1) Emergency use of the entry battery helped prevent potential
loss of dc-main bus A, which could have led to loss of com-
munications between spacecraft and ground and other vital CM
functions.

(2) Available emergency powerdown lists facilitated rapid re-
duction 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 in-
strumentation 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.

b. When these procedures failed to arrest the oxygen loss
Mission Control directed the crew to shut down fuel cells 1
and 3 by closing the hydrogen and oxygen flow valves.

Determinations

(1) 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. 1.

Lunar Module Activation

42. Findings

a. With imminent loss of oxygen from oxygen tanks no. 1 and
no. 2, and failing electrical power in the CM, it was
necessary to use the lunar module (LM) as a "lifeboat" for
the return to Earth.

b. 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 situa-
tion. 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 check-
list procedure which manually transferred the CM alignment to
the LM.

Determinations

(1) 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 modifi-
cations to the checklists which sufficiently shortened the
time normally required for powering up the LM.

43. Findings

a. During the LM powerup and the CSM powerdown, there was a brief time
interval during which Mission Control gave the crew directions which
resulted in neither module having an active attitude control system. 

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

c. The Command Module Pilot (CMP) stated that he was not con-
cerned 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 re-
establish 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.


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 re-
duction of LM power for use in case an LM battery problem
developed. Procedures for use of CM water in the LM also
were developed for use if needed.

e. Toward the end of the mission, sufficient consumable margins
existed to allow usage rates to be increased above earlier
planned levels. This was done.

f. When 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

(1) 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.


Modification of LM Carbon Dioxide Removal System

46. Findings 

a. The lithium hydroxide (LiOH) cartridges, which remove water
and carbon dioxide from the LM cabin atmosphere, would have
become ineffective due to saturation at about 100 hours.

b. Mission rules set maximum allowable carbon dioxide partial
pressure at 7.5mm Hg. LiOH cartridges are normally changed
before cabin atmosphere carbon dioxide partial pressure
reaches this value.

c. Manned Spacecraft Center engineers devised and checked out a
procedure for using the CM LiOH canisters to achieve carbon
dioxide removal. Instructions were given on how to build a
modified cartridge container using materials in the space-
craft.

d. The crew made the modification at 93 hours, and carbon
dioxide partial pressure in the LM dropped rapidly from
7.5mm Hg to 0.lmm Hg. 

e. Mission Control gave the crew further instructions for
attaching additional cartridges in series with the first
modification. After this addition, the carbon dioxide partial
pressure remained below 2mm Hg for the remainder of the Earth-
return 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.

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 1 and 2 are located. This venting con-
tinued 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 1 and 2 supplied 9
of the 11 amperes total. By 97:23 the load-sharing had re-
turned 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

(1) 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.

(3) 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.

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

c. Mission Control worked out a procedure for using LM battery
power to recharge CM batteries A and B. This procedure used
the electrical umbilical between the LM and the CM which
normally carried electrical energy from the CM to the LM.
Re procedure was nonstandard and was not included in check-
lists

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 the 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 rturn 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.

A 3.2-fps correction at 137:40 using LM RCS thrusters, to
assure that the CM would reenter the Earth's atmosphere at
the center of its corridor.

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 param-
eters.

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 splash-
down where the best recovery forces were located; it shortened
the flight time to increase safety margins in the use of elec-
trical 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 be-
cause 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.

e. The reentry preparation time had to be extended in order to
accomplish the additional steps required by the unusual situa-
tion.

f. In order to conserve the CM batteries, LM jettison was de-
layed 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

(1) The procedures developed worked well and generated no new
hazards beyond those unavoidably inherent in using procedures
which have not been carefully developed, simulated, and
practiced over a long training period.

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

51. Findings

a. During the reentry preparations, after SM jettison, there was
a half-hour period of very poor communications with the CM
due to the spacecraft being in a poor attitude with the LM
present.

b. This condition was not recognized by the crew or by Mission
Control.

Determination

Some of the reentry preparations were unnecessarily prolonged by
the poor communications, but since the reentry preparation time-
line 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 occur-
rence of gimbal lock in itself would not have significantly in-
creased the crew hazard.


PART 4.  RECOMMENDATIONS

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

     a. Remove from contact with the oxygen all wiring, and the unsealed
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 com-
bustle 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 po-
tential 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 out-
of-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 EM and the CM should be re-
viewed to determine whether steps should be taken to enhance their po-
tential for use in a "lifeboat" mode. 

5. The Manned Spacecraft Center should complete the special tests and
analyses now underway in order to understand more completely the details
of the Apollo 13 accident. In addition, the lunar module power system
anomalies should receive careful attention. Other NASA Centers should
continue their support to MSC in the areas of analysis and test. 

6. Whenever significant anomalies occur in critical subsystems during
final preparation for launch, standard procedures should require a presen-
tation of all prior anomalies on that particular piece of equipment, in-
cluding those which have previously been corrected or explained.  Further-
more, critical decisions involving the flightworthiness of subsystems
should require the presence and full participation of an expert who is
intimately familiar with the details of that subsystem. 

7. NASA should conduct a thorough reexamination of all of its spacecraft,
launch vehicle, and ground systems which contain high-density oxygen, or
other strong oxidizers, to identify and evaluate potential combustion
hazards in the light of information developed in this investigation. 

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

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