


APOLLO/SATURN V SPACE VEHICLE
-----------------------------

The primary flight hardware of the Apollo Program consists of the
Saturn V launch vehicle and Apollo spacecraft.
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 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 electrically with the S-II
stage. It also interfaces structurally, electrically, and
pneumatically with ground support equipment (GSE) through two
umbilical service arms, three tail service masts, and certain
electronic systems by antennas. The S-IC stage is instrumented
for operational measurements or signals which are transmitted by
its independent telemetry system.

S-II Stage

The S-II stage 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 inter-stage 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 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 inter-stage weight of 8,100 pounds is
not included in the stated weights. The stage is instrumented for
functional measurements or signals which are transmitted by its
independent telemetry system.

The high performance J-2 engine as installed in the S-IVB stage
has a multiple start capability. The S-IVB J-2 engine is
scheduled to produce a thrust of 203,000 pounds during its first
burn to earth orbit and a thrust of 178,000 pounds (mixture mass
ratio of 4.5:1) during the first 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
4,310 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 is a conical structure which provides a
structural load path between the LV and SM and also supports the
LM.  Aerodynamically, the SLA smoothly encloses the irregularly
shaped LM and transitions the space vehicle diameter from that of
the upper stage of the LV to that of the SM. The SLA also
encloses the nozzle of the SM engine and the high gain antenna.

Spring thrusters are used to separate the LM from the SLA. After
the CSM has docked with the LM, mild charges are fired to release
the four adapters-which secure the LM in the SLA. Simultaneously,
four spring thrusters mounted on the lower (fixed) SLA panels
push against the LM landing gear truss assembly to separate the
spacecraft from the launch vehicle.


Service Module
--------------

The service module (SM) provides the main spacecraft
propulsion and maneuvering capability during a mission. The SM
provides most of the spacecraft consumables (oxygen, water,
propellant, and hydrogen) and supplements environmental,
electrical power, and propulsion requirements of the CM. The SM
remains attached to the CM until it is jettisoned just before CM
atmospheric entry.

Structure.

        The basic structural components are forward and aft
(upper and lower) bulkheads, six radial beams, four sector
honeycomb panels, four reaction control system honeycomb panels,
aft heat shield, and a fairing.  The forward and aft bulkheads
cover the top and bottom of the SM.  Radial beam trusses
extending above the forward bulkhead support and secure the CM.
The radial beams are made of a solid aluminum alloy which has been
machined and chem-milled to thicknesses varying between 2 inches
and ,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 sub-system 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 radiator and an RCS
engine quad on its exterior and contains the SPS engine fuel
storage tank which feeds the fuel sump tank in sector five. The
center section contains two helium tanks and the SPS engine. The
tanks are used to provide helium pressurant for the SPS
propellant tanks.

Propulsion.

         Main spacecraft propulsion is provided by the 20,500
pound thrust SPS. The SPS engine is a restartable, non-
throttleable engine which uses nitrogen tetroxide (N2O4) 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 mid-course corrections, lunar
orbit insertion, trans-earth 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) 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 an aluminum sandwich construction
consisting of a welded aluminum inner skin, a bonded aluminum
honeycomb core, and an 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)
has been arranged to provide for the expected duties of crew
members. These duties fall into the categories of Commander, CM
Pilot, and LM Pilot, occupying the left, center, and right
couches, respectively. The CM Pilot also acts as the principal
navigator. All controls have been designed so they can be
operated by astronauts wearing gloves. The controls are
predominantly of four basic types: toggle switches, rotary
switches with click-stops, thumb-wheels, and push buttons.
Critical switches are guarded so that they cannot be thrown
inadvertently.  In addition, some critical controls have locks
that must be released before they can be operated.

Flight controls are located on the left center and left side of
the MDC, opposite the Commander. These include controls for such
subsystems as stabilization and control, propulsion, crew safety,
earth landing, and emergency detection. One of two guidance and
navigation computer panels also is located here, as are velocity,
attitude, and altitude indicators.

The CM Pilot faces the center of the console, and thus can reach
many of the flight controls, as well as the system controls on
the right side of the console. Displays and controls directly
opposite him include reaction control, propellant management,
caution and warning, environmental control, and cryogenic
storage systems. The rotation and translation controllers used
for attitude, thrust vector, and translation maneuvers are
located on the arms of two crew couches. In addition, a rotation
controller can be mounted at the navigation position in the lower
equipment bay.

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


Lunar Module
------------

The lunar module (LM) 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 1,050 pounds of thrust to 6,300
pounds. Throttle positions above this value automatically produce
full thrust to reduce combustion chamber erosion. Nominal full
thrust is 9,870 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 3,500 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 2,008 pounds, of which 71 pounds are
unusable. Oxidizer weight is 3,170 pounds, of which 92 pounds are
unusable. The APS has a limit of 35 starts, must have a
propellant bulk temperature between 50 F. and 90 F. prior to
start, must not exceed 460 seconds of burn time, and has a system
life of 24 hours after pressurization.

Electrical power system.

           The electrical power system (EPS) contains
six batteries which supply the electrical power requirements
of the LM during un-docked 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: pre-launch
checkout and launch operations; tracking the space vehicle to
determine its present and future positions; securing information
on the status of the flight crew and space vehicle systems (via
telemetry); evaluation of telemetry information; commanding the
space vehicle by transmitting real-time and updata commands to
the onboard computer; and voice communication between flight
and ground crews.

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

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

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

b. Kennedy Space Center (KSC), Cape Kennedy, Florida. The space
vehicle is launched from KSC and controlled from the Launch
Control Center (LCC). 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.


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. 

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

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

a. Flight crew observations

b. Flight controller real-time observations

c. Review of telemetry data received from tape recorder playback

d. Trend analysis of actual and predicted values

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




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.



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. EST,
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-IC 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 trans-lunar 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 trans-lunar
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.

Trans-lunar 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 the lunar surface
at 6 53' S., 30 53' W. versus the targeted 3 S., 30 W. Therefore,
the second S-IVB APS (trim) burn was not required. 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
mid-course 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 circum-lunar trajectory with a predicted closest approach
to the moon of 62 n. mi. All SPS burn parameters were normal. The
accuracy of MCC-2 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. EST), the crew reported
an undervoltage alarm on the CSM main bus B. Pressure was rapidly
lost in SM oxygen tank no. 2 and the current in fuel cells 1 and 3
dropped to zero due to loss of their oxygen supply. At this point,
the 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 mid-course 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.

Trans-earth Coast

At pericynthion plus 2 hours ("PC+2", 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 delta V
(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 (firing the RCS engine quad clusters to put the
CSM/LM into a slow roll, to allow incident solar energy to fall on
the entire surface of both CSM/LM, to reduce the load on the ECS)
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 14
second burn (at 10% throttle) produced a delta V 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
conditions of the CM, with first reported receipt of S-band
signal at 101:53 g.e.t. Thermal conditions on all CSM systems
observed appeared to be in order for entry.

Due to the unusual spacecraft configuration, new procedures
leading to entry were developed and verified in ground-based
simulations.  The resulting timeline called for a final mid-course
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 2,000 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. EST, 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, the helicopter carrier USS IWO JIMA. The
crew, picked up by a recovery helicopter, was safe aboard the
ship at 1:53 p.m. EST, less than an hour after landing.





REVIEW AND ANALYSIS OF APOLLO 13 ACCIDENT
-----------------------------------------



PART 1. INTRODUCTION

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

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



OXYGEN TANK NO. 2 HISTORY


DESIGN
------

On February 26, 1966, the North American Aviation Corporation,
now Rockwell International, 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.

Each oxygen tank has an outer shell and an inner shell, arranged
to provide a vacuum space to reduce heat leak, and a dome enclosing
paths into the tank for transmission of fluids and electrical power
and signals. The space between the shells and the space in the dome
are filled with insulating materials. Mounted in the tank are two
tubular assemblies. One, called the heater tube, contains two
thermostatically protected heater coils and two small fans driven
by 1,800 rpm motors to stir the tank contents. The other, called
the quantity probe, consists of an upper section which supports
a cylindrical capacitance gauge 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 gauge. In addition, a temperature sensor is mounted
on the outside of the quantity probe near the head.  Wiring for
the gauge, 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 fill line from the exterior of the SM enters the oxygen tank
and connects to the inner cylinder of the capacitance gauge 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 gauge cylinder.
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 then, 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
gauge in the supply line and a pressure switch near this gauge
is provided to turn on the heaters in the oxygen tank if the
pressure drops below a pre-selected 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 "super-critical" means that the
oxygen is maintained at a temperature and pressure which assures
that it is in a homogeneous, single-phase "fluid" state.)

The burst pressure of the oxygen tank is about 2,200 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 1,000 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 OXYGEN TANK & ITS SHELF
---------------------------

Two oxygen tanks are mounted on a shelf in bay 4 of the SM. The
bottom of the oxygen shelf houses some of the oxygen system
instrumentation and wiring, largely covered by insulation.


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.
Re-welding 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 vacuum pumps. The tank was then leak-tested
at 500 psi and proof tested at 1,335 psi with helium.

After the helium proof test, the tank was filled with liquid
oxygen and pressurized to a proof pressure of 1,335 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 by applying warm gas at about 30 psi to the vent
line to force the liquid oxygen (LOX) in the tank out the fill
line. No difficulties were recorded in this de-tanking 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 Rockwell on May 3, 1967.


ASSEMBLY AND TEST AT 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
standard proof-pressure, leak, and functional checks. One valve
on the shelf leaked and was repaired, but no anomalies were noted
with regard to oxygen tank no. 2, and therefore no re-work of
oxygen tank no. 2 was required. None of the oxygen tank testing
at Rockwell required 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 vacuum-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.

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 the bay. 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 two
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 close-out 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 two
inches. It now seems likely that the shelf was first accelerated
upward and then dropped.

The remaining bolt was then removed, the incident recorded, and
the oxygen shelf was removed without further difficulty.
Following removal, the oxygen shelf was re-tested to check shelf
integrity, including proof-pressure tests, leak tests, and
functional tests of pressure transducers and switches, thermal
switches, and vacuum-ion pumps. Cryogenic testing was conducted.
Visual inspection revealed no problems. 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 only a two inch drop. 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 a pressure of 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 per cent 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 de-tanking 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, de-tanking operations were resumed,
after discussions of the problem had been held, with KSC, MSC,
Rockwell, 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 per
cent. Further discussions between KSC, MSC, Rockwell, and Beech
personnel considered that the problem might be due to a leak in
the path between the fill line and the quantity probe due to loose
fit in the sleeves and tube. Such a leak would allow the gaseous
oxygen (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" de-tanking 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 65V DC GSE power supply, and,
about 1 hours later, the fans were turned on to add more heat and
mixing. After 6 hours of heater operation, the quantity had only
decreased to 35 per cent, 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 per cent 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 Rockwell 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 gauge 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 gauge. 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 per cent 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 de-tanking
problem was considered by the Apollo organization. At this point,
the "shelf drop" incident on October 21, 1968, at Rockwell was not
considered and it was felt that the apparently normal de-tanking
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, Rockwell, Beech, and NASA
headquarters, emphasis was directed toward the possibility and
consequences of a loose fill tube, while very little attention
was paid to the extended operation of heaters and fans, except
to note that they apparently operated during and after the
de-tanking 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, 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 65V DC supply similar to the power used at
KSC during the de-tanking sequence.

Subsequent investigations have shown that the thermostatic
switches used, while rated as satisfactory for the 28V DC
spacecraft power supply, could not open properly at 65V 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 de-tanking 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 1,000 F during the de-tanking. 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 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 installation on the lunar surface near the lunar
module (LM) landing point: (1) a lunar passive seismometer to
measure and relay meteoroid impact and moon quakes, 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
gauge 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, back-up 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 crew Commander James A. Lovell and prime crew Lunar Module
Pilot Fred Haise were immune to rubella, but that prime crew
Command Module Pilot Thomas K. ("Ken") Mattingly III did not have
immunity. Consequently, following two days of intensive simulator
training at KSC, back-up Command Module Pilot John L. ("Jack")
Swigert, Jr., was substituted in the prime crew to replace
Mattingly. Swigert had trained for several months with the
back-up 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. EST, 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-IC second stage.  As a result, the remaining four S-IC
engines burned 34 seconds longer than planned and the S-IVB third
stage burned a few seconds longer than planned. At orbital
insertion, the velocity was within 1.2 feet per second of
the planned velocity. Moreover, an adequate propellant margin
was maintained in the S-IVB for the trans-lunar injection burn.

Orbital insertion was at 00:12:39 ground elapsed time (g.e.t.).
The initial one and a half earth orbits before trans-lunar
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 trans-lunar
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 mid-course correction maneuver at
11:41:23 g.e.t. was not necessary.  After TLI, Apollo 13 was
calculated to be on a free-return trajectory with a predicted
closest approach to the lunar surface of 210 nautical miles.

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 provide propulsive energy.  On Apollo 13, the
plan was to impact the S-IVB stage on the lunar surface in
proximity to the seismometer placed in the Ocean of Storms
by the crew of Apollo 12.

Two hours after TLI, the S-IVB attitude thrusters were ground-
commanded on to adjust the stage's trajectory toward the
designated impact at latitude 3 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 trans-lunar
and trans-earth 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 mid-course 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 mid-course correction
maneuver was a 23.2 feet per second hybrid transfer burn which
took Apollo 13 off a 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 free-
return trajectories until lunar orbital 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
performance of oxygen tank no. 2 was normal, so far as
telemetered data and crew observations indicate. At 46:40:02, the
crew turned on the fans in oxygen tank no. 2 as a routine
operation. Within 3 seconds, the oxygen tank no. 2 quantity
indication changed from a normal reading of about 82 percent full
to an obviously incorrect reading "off-scale high," or over 100
per cent. Analysis of the electrical wiring of the quantity gauge
shows that this erroneous reading could be caused by either a
short circuit or an open circuit in the gauge wiring or a short
circuit between the gauge plates. Subsequent events indicated
that a short circuit 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
gauge 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 in-flight inspection of the
LM. Ground tests before launch had indicated the possibility of a
high heat-leak rate in the LM descent stage super-critical helium
tank. Crew verification of actual pressures found the helium
pressure to be within normal limits. (Super-critical 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 pre-sleep 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 93 seconds 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 turn-on 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 the effects of oxygen escaping from the
tank caused the removal of the panel covering bay 4, and a
relatively slow leak in oxygen tank no. 1 or its lines or valves.

Photos of the SM taken by the crew later in the mission show
the panel missing, the fuel cells on the shelf above the oxygen
shelf tilted, and the high-gain antenna damaged.

The resultant loss of oxygen made the fuel cells inoperative,
leaving the CM with batteries normally used only during re-entry
as the sole power source, and with only that oxygen contained in a
surge tank and re-pressurization packages (used to re-pressurize
the CM after cabin venting). The LM, therefore, became the only
source of sufficient electrical power and oxygen to permit safe
return of the crew to Earth.




DETAILED CHRONOLOGY FROM BEFORE ACCIDENT TO 5 MINUTES AFTER
-----------------------------------------------------------


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 disturbance
              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
              1,008.3 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 following
              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 temperature
              begins a 1.66 F. increase in 36 seconds.

55:55:01      Oxygen flow rates to fuel cells 1 and 3 approaches zero
              after decreasing for 7 seconds.

55:55:02      The surface temperature of the service module oxidizer 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 bus B 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 per cent. 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 main 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 main 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. 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.




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 by or for the Board after the
accident.

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 cer-
tainty, the nature of the occurrence is sufficiently
understood to per mit 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
........

In evaluating telemetry data, consideration must be given to the
fact that the Apollo pulse code modulation (PCM) system samples
data in time and quantizes in amplitude.

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

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

55:53:22.772  11.1 ampere "spike" recorded in fuel cell 3 current
              followed by drop in current and rise in voltage
              typical of removal of power from one fan motor,
              indicating opening of motor circuit.

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

The evidence points strongly to an electrical short circuit with
arcing as the initiating event. About 2.7 seconds after the fans
were turned on in the SM oxygen tanks, an 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 gauge 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 gauge in oxygen tank no.
2 had failed at 46:40 g.e.t. There is no evidence tying the
quantity gauge failure directly to accident initiation,
particularly in view of the very low energy available from the
gauge.)

This likelihood of electrical initiation is enhanced by the high
probability that the electrical wires within the tank were
damaged during the abnormal de-tanking 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 to 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
              typical of one fan motor -- indicating opening of
              another motor circuit.

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

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

55:54:30      Oxygen tank no. 2 quantity gauge reading drops from
              full scale (to which it had failed at 46:40 g.e.t.)
              to zero and then read 75 per cent full. This behavior
              indicates the gauge short circuit may have corrected
              itself.

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

55:54:45      Oxygen tank no. 2 pressure reading reaches maximum
              recorded value of 1008 psia.

55:54:52.763  Oxygen tank no. 2 pressure reading had dropped
              to 996 psia.

The available evidence points to a combustion process as the
cause of the pressure and temperature increases recorded in
oxygen tank no. 2. The pressure reading for oxygen tank no. 2
began to increase about 13 seconds after the first electrical
spike, and about 55 seconds later the temperature began to
increase. The temperature sensor reads local temperature, which
need not represent bulk fluid temperature. Since the rate of
pressure rise in the tank indicates a relatively slow propagation
of burning, it is likely that the region immediately around the
temperature sensor did not become heated until this time.

There are materials within the tank that can, if ignited in the
presence of super-critical 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
super-critical 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 determined; it is believed
to be associated with a change in reaction rate as combustion
proceeded through various Teflon elements.

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



LOSS OF OXYGEN TANK NO. 2 SYSTEM 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.

After the relatively slow propagation process described above
took place, there was a relatively abrupt loss of oxygen tank no.
2 integrity. About 69 seconds after the pressure began to rise,
it reached the peak recorded (1008 psia), the pressure at which
the cryogenic oxygen tank relief valve is designed to be fully
open. Pressure began a decrease for 8 seconds, dropping to 996
psia before readings were lost.

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

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 re-acquired, 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 re-acquisition 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.

The data are not adequate to determine precisely the way in which
oxygen tank no. 2 system lost its integrity. However, available
information, analyses, and tests performed during this
investigation indicate that most probably the combustion within
the pressure vessel ultimately led to localized heating and
failure at the pressure vessel closure. It is at this point, the
upper end of the quantity probe, that the 0.5 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, 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-    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 oxygen tank no. 1 had lost its integrity. Pressure de-
creases 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.




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 blow-off 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 un-powered.
The oxygen tank no. 2 temperature and quantity gauges 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 DC 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 minutes. He also reported fluctuations in the oxygen
oxygen tank no. 2 quantity, followed by a return to the
off-scale high position.

When fuel cells 1 and 3 electrical output readings went to zero,
the ground controllers could not be certain that the cells had
not somehow been disconnected from their respective busses and
were not otherwise all right. Attention continued to be focused
on electrical problems.

Five minutes after the accident, controllers asked the crew to
connect fuel cell 3 to DC main bus B in order to be sure that the
configuration was known. When it was realized that fuel cells 1
and 3 were not functioning, the crew was directed to perform an
emergency power-down to lower the load on the remaining fuel cell.
Observing the rapid decay in oxygen tank no. 1 pressure,
controllers asked the crew to switch power to the oxygen tank no.
2 instrumentation. When this was done, and it was realized
that oxygen tank no. 2 had failed, the extreme seriousness of the
situation became clear.

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

It was obvious by about 1 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 all 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.






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


                       Required:          85 hours

                       Available in LM:   53 hours

                       Available in CM:  182 hours





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 super-critical oxygen.

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

   e. During the special de-tanking 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 65V DC in order to protect the heaters from over-
      heating. The switches were only rated at 30V 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 90 seconds immediately prior to the
      accident.


2. Determinations

   a. The cause of the failure of oxygen tank no. 2 was combustion
      within the tank.

   b. Analysis showed that the electrical energy flowing into the
      tank could not account for the observed increases in pressure
      and temperature.

   c. The heater, temperature sensor, and quantity probe did not
      initiate the accident sequence.

   d. 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.

   e. The protective thermostatic switches on the heaters in oxygen
      tank no. 2 failed closed during the initial portion of the first
      special de-tanking operation. This subjected the wiring in the
      vicinity of the heaters to very high temperatures which have
      been subsequently shown to severely degrade Teflon insulation.

   f. 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 1 ampere fuses
      on the fan motors would pass sufficient energy to ignite the
      insulation by the mechanism of an electric arc.

   g. The combustion of Teflon wire insulation alone could release
      sufficient heat to account for the observed increases in tank
      pressure and local temperature, and could locally overheat and
      fail the tank or its associated tubing. The possibiIity of such
      failure at the top of the tank was demonstrated by subsequent
      tests.

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




SECONDARY EFFECTS OF TANK FAILURE

1. Findings

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

      (1) A "bang" heard by the crew,

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

      (3) Momentary loss of telemetry,

      (4) Closing of several valves by shock loading,

      (5) Loss of integrity of the oxygen tank no. 1 system,
      
      (6) Slight temperature increases in bay 4 and adjacent sectors of
          the SM,

      (7) Loss of the panel covering bay 4 of the SM, as observed and
          photographed by the crew,

      (8) Displacement of the fuel cells as photographed by the crew, and

      (9) 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 inter-connected 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.



2. Determinations

   a. 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.

   b. 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 subsequent tests.

   c. 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.

   d. Photographs of the SM by the crew did not establish the
      condition of oxygen tank no. 2.

   e. The high-gain antenna damage probably resulted from striking
      by the panel, or a portion thereof, as it left the SM.

   f. The loss of pressure on oxygen tank no. 1 and the subsequent
      loss of power resulted from the tank no. 2 failure.

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

   h. 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.








