NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
RELEASE NO: 70-50K  April 2, 1970
PROJECT: APOLLO 13
GENERAL RELEASE
APOLLO 13 COUNTDOWN
  Lightning Precautions
  May Launch Opportunities
  Apollo 13 Flight Profile
LAUNCH, MISSION TRAJECTORY AND MANEUVER DESCRIPTION
  Launch
  Launch Events
  Apollo 13 Mission Events
  Earth Parking Orbit (EPO)
  Translunar Injection (TLI)
  Transposition, Docking, and Ejection (TD&E)
  Saturn Third Stage Lunar Impact 24-26
  Translunar Coast
  Lunar Orbit Insertion (LOI)
  Descent Orbit Insertion (DOI)
  Lunar Module Separation
  CSM Circularization
  Power Descent Initiation (PDI), Lunar Landing
  Lunar Surface Exploration
  Ascent, Lunar Orbit Rendezvous
  Ascent Stage Deorbit
  Transearth Injection (TEI)
  Transearth Coast
  Recovery Operations
APOLLO 13 ONBOARD TELEVISION
  Apollo 13 TV Schedule
APOLLO 13 SCIENCE
  Lunar Orbital Photography
  Charged Particle Lunar Environment Experiment (CPLEE)
  Lunar Atmosphere Detector (LAD)
  Lunar Heat Flow Experiment (HFE)
  Passive Seismic Experiment (PSE)
  Solar Wind Composition
  Dust Detector
  Field Geology Investigations
  SNAP-27
PHOTOGRAPHIC EQUIPMENT
LUNAR DESCRIPTION
  Physical Facts
  Landing Site
APOLLO 13 FLAGS, LUNAR MODULE PLAQUE
SATURN V LAUNCH VEHICLE
  First Stage
  Second Stage
  Third Stage
  Instrument Unit
  Propulsion
COMMAND AND SERVICE MODULE STRUCTURE, SYSTEMS
  Launch Escape System (LES)
  Command Module (CM) Structure
  Service Module (SM) Structure
  Spacecraft-LM Adapter (SLA) Structure
  CSM Systems
LUNAR MODULE STRUCTURES, WEIGHT
  Ascent Stage
  Descent Stage
  Lunar Module Systems
APOLLO 13 CREW AND CREW EQUIPMENT
  Life Support Equipment-Space Suits
  Apollo Lunar Hand Tools
  Apollo 13 Crew Menu
  Personal Hygiene
  Medical Kit
  Survival Gear
  Biomedical Inflight Monitoring
  Training
  Crew Biographies
LAUNCH COMPLEX 39
MISSION CONTROL CENTER
MANNED SPACE FLIGHT NETWORK
  Network Computers
  Apollo Ship Vanguard
  Apollo Range Instrumentation Aircraft (ARIA)
CONTAMINATION CONTROL PROGRAM
LUNAR RECEIVING LABORATORY (LRL)
LUNAR RECEIVING LABORATORY TENTATIVE SCHEDULE
SCHEDULE FOR TRANSPORT OF SAMPLES, SPACECRAFT AND CREW
APOLLO PROGRAM MANAGEMENT
  Apollo/Saturn Officials
  Major Apollo/Saturn V Contractors
FOR RELEASE: THURSDAY A.M.
April 2, 1970
RELEASE NO: 70-50
APOLLO 13 THIRD LUNAR LANDING MISSION
Apollo 13, the third U.S. manned lunar landing mission, will be launched
April 11 from Kennedy Space Center, Fla., to explore a hilly upland region
of the Moon and bring back rocks perhaps five billion years old.
The Apollo 13 lunar module will stay on the Moon more than 33 hours and
the landing crew will leave the spacecraft twice to emplace scientific
experiments on the lunar surface and to continue geological
investigations. The Apollo 13 landing site is in the Fra Mauro uplands;
the two National Aeronautics and Space Administration previous landings
were in mare or "sea" areas, Apollo 11 in the Sea of Tranquillity and
Apollo 12 in the Ocean of Storms.
Apollo 13 crewmen are commander James A. Lovell, Jr.; command module pilot
Thomas K. Mattingly III, and lunar-module pilot Fred W. Haise, Jr.  Lovell
is a U.S. Navy captain, Mattingly a Navy lieutenant commander, and Haise a
civilian.
3/26/70
Launch vehicle is a Saturn V. Apollo 13 objectives are
* Perform selenological inspection, survey and sampling of materials in a
preselected region of the Fra Mauro formation.
* Deploy and activate an Apollo Lunar Surface Experiment Package (ALSEP).
* Develop man's capability to work in the lunar environment
* Obtain photographs of candidate exploration sites.
Currently 11 television transmissions in color are scheduled: one in Earth
orbit an hour and a half after launch, three on the outward voyage to the
Moon; one of the landing site from about nine miles up;  two from the
lunar surface while the astronauts work outside the spacecraft; one at
the command service module/lunar module docking operation; one of the Moon
from lunar orbit; and two on the return trip.
The Apollo 13 landing site is in the hilly uplands to the north of the
crater Fra Mauro. Lunar coordinates for the landing site are 3.6 degrees
south latitude by 17.5 degrees west longitude, about 95.6 nautical Miles
east of the Apollo 12 landing point at Surveyor III crater.
Experiments emplaced at the Fra Mauro site as part of the ALSEP III will
gather and relay long-term scientific data to Earth for at least a- year
on the Moon's physical and environmental properties. Five experiments
are contained in the ALSEP: a lunar passive seismometer will measure and
relay meteoroid impacts and moonquakes, a heat flow experiment will
measure the heat flux from the lunar interior to the surface and
conductivity of the surface materials to a depth of about. 10 feet; a
charged particle lunar environment experiment will measure protons and
electrons to determine the effect of the solar wind on the lunar
environment; a cold cathode gauge experiment will measure density and
temperature variations in the lunar atmosphere; and a dust detector
experiment.
The empty third stage of the Saturn V launch vehicle will be targeted to
strike the Moon before the lunar landing and its impact will be recorded
by the seismometer left by the Apollo 12 astronauts last November. The
spent lunar module ascent stage, as in Apollo 12, will be directed to
impact the Moon after rendezvous and final ML separation to provide a
signal to both seismometers.
Candidate future Apollo landing sites -- Censorinus, Davy Rille, and
Descartes -- will be photographed with a large-format lunar topographic
camera carried for the first time on Apollo 13. The lunar topographic
camera will make high- resolution 4.5 inch square black-and-white photos
in overlapping sequence for mosaics or in single frames. The camera mounts
in the command module crew access hatch window when in use. After lunar
orbit rendezvous with the lunar module and ML jettison the command module
will make a plane-change maneuver to drive the orbital track over
Descartes and Davy Rille for topographic photography.
The Apollo 13 light profile in general follows those flown by Apollos 11
and 12 with one major exception: lunar orbit insertion burn no.  2 has
been combined with descent orbit insertion and the docked spacecraft
will be placed into a 7x57 nautical mile lunar orbit by use of the service
propulsion system. Lunar module descent propellant is conserved by
combining these maneuvers to provide 15 seconds Or additional hover time
during the landing.
Lunar surface touchdown is scheduled to take place at 9:55 p.m.  EST April
15, and two periods of extravehicular activity are planned at 2:13 a.m.
EST April 16 and 9:58 p.m. EST April 16. The LM ascent stage will lift off
at 7:22 a.m. April 17 to rejoin the orbiting command module after more
than 33 hours on the lunar surface. Apollo 13 will leave lunar orbit at
1:42 p.m. EST April 18 for return to Earth. Splashdown in the mid-Pacific
just south of the Equator will be at 3:17 p.m. EST April 21.
After the spacecraft has landed, the crew will put on clean coveralls and
filter masks passed in to them through the hatch by a swimmer, and then
transfer by helicopter to a Mobile Quarantine Facility (MQF) on the USS
Iwo Jima. Them MQF and crew will be offloaded in Hawaii and placed aboard a
C-141 aircraft for the flight back to the Lunar Receiving Laboratory at
the Manned Spacecraft Center in Houston.  The crew will remain in
quarantine up to 21 days from completion of the second EVA.
The crew of Apollo 13 selected the call signs Odyssey for the command
service module and Aquarius for the lunar module. When all three crewmen
are aboard the command module, the call sign will be "Apollo 13." As in
the two previous lunar landing missions, an American flag will be emplaced
on the lunar surface. A plaque bearing the date of the Apollo 13 landing
and the crew signatures is attached to the LM.
Apollo 13 backup crewmen are USN commander John W. Young, commander
civilian John L. Swigert, Jr., command module pilot and USAF Major Charles
M. Duke, Jr., lunar module pilot. APOLLO 13 COUNTDOWN
Precount activities for the Apollo 13 launch begin about T-4 days, when
the space vehicle will be prepared for the start of the Official
countdown. During precount, final space vehicle ordnance installation and
electrical connections will be accomplished.  Spacecraft gaseous oxygen
and gaseous helium systems will be serviced, spacecraft batteries will be
installed, and LM and CSM mechanical buildup will be completed. The CSM
fuel cells will be activated and CSM cryogenics (liquid oxygen - liquid
hydrogen) will be loaded and pressurized.
The countdown for Apollo 13 will begin at T-28 hours and will continue to
T-9 hours, at which time a built-in hold is planned prior to the start of
launch vehicle propellant loading.
Following are some of the major operations in the final count:
T-28 hours  			Official countdown starts
				LM stowage and cabin closeout
  				(T-31: 30 to T-18: 00 )
T-27 hours, 30 minutes  	Install and connect LV flight
 				batteries ( to T-23 hours )
T-22 hours, 30 minutes  	Topoff of LM super critical helium
  				(to T-20 hours, 30 minutes )
T-l9 hours, 30 minutes 		LM The thermal shield installation
  				(to T-15 hours, 30 minutes) CSM
  				crew stowage (T-l9 to T-12 hours,
  				30 minutes)
T-16 hours 			LV range safety checks (to T-15
 				hours)
T-15 hours  			Installation of ALSEP FCA to
  				T-14 hours, 45 minutes)
T-11 hours, 30 minutes  	Connect LV safe and arm devices
  				(to 10 hours, 5 minutes) CSM pre-
 				ingress operations (to T-8 hours
  				45 minutes)
T-10 hours, 15 minutes  	Start MSS move to parksite
T-9 hours 			Built-in hold for 9 hours and 13
 				minutes. At end of hold, pad is
 				-cleared for LV propellant loading
T-8 hours, O5 minutes 		Launch vehicle propellant loading - Three
				stages (LOX in first stage, LOX and LH2 in
				second and third stages).  Continues
				thru T-3 hours 38 minutes
T-4 hours, 17 minutes 		Flight crew alerted
T-4 hours, 02 minutes 		Medical examination
T-3 hours, 32 minutes 		Breakfast
T-3 hours, 30 minutes 		One-hour hold
T-3 hours, 07 minutes 		Depart Manned Spacecraft Operations
				Building or LC-39 via crew transfer
				van.
T-2 hours, 55 minutes 		Arrive at LC-39
T-2 hours, 40 minutes 		Start flight crew ingress
T-2 hours 			Mission Control Center - Houston/
				spacecraft command checks
T-l hour, 55 minutes 		Abort advisory system checks
T-l hour, 51 minutes		Space Vehicle Emergency Detection
				System (EDS) test
T-43 minutes 			Retract Apollo access arm to stand-
				by position (12 degrees)
T-42 minutes 			Arm launch escape system
T-40 minutes 			Final launch vehicle range safety
				checks (to 35 minutes)
T-30 minutes 			Launch vehicle power transfer test
				LM switch over to internal power
T-20 minutes to T-10 minutes 	Shutdown LM operational instrumentation
T-15 minutes  			Spacecraft to internal power
T-6 minutes 			Space vehicle final status checks
T-5 minutes, 30 seconds 	Arm destruct system
T-S minutes 			Apollo access arm fully retracted
T-3 minutes, 7 seconds 		Firing command (-automatic
				sequence)
T-50 seconds 			Launch vehicle transfer to internal
 				power
T-8.9 seconds 			Ignition sequence start
T-2 seconds 			All engines running
T-0 				Liftoff
Note: Some changes in the above countdown are possible as a result of
experience gained in the countdown demonstration which occurs about 10
days before launch.
Lightning Precautions
During the Apollo 12 mission the space vehicle was subjected to two
distinct electrical discharge events. However, no serious damage occurred
and the mission proceeded to a successful conclusion. Intensive
investigation led to the conclusion that no hardware changes were necessary
to protect the space vehicle from-similar events. For Apollo 13 the
mission rules have been revised to reduce the probability that the space
vehicle will be launched into cloud formations that contain conditions
conducive to initiate similar electrical discharges although flights into
all clouds is not precluded.
May Launch Opportunities
The three opportunities established for May -- in case the launch is
postponed from April 11 -- provide, in effect, the flexibility of a choice
of two launch attempts.  The optimum May launch window occurs on May 10.
The three day window permits a choice of attempting a launch 24 hours
earlier than the optimum window and if necessary a further choice of a 24
hour or 48 hour recycle. It also permits a choice of making the first
launch attempt on the optimum day with a 24-hour recycle capability. The
May 9 window (T-24 hrs) requires an additional 24 hours in lunar orbit
before initiating powered descent to arrive at the landing site at the
same time and hence have the same Sun angle for landing as on May 10.
Should the May 9 window launch attempt be scrubbed, a decision will be
made at that time, based on the reason for the scrub, status of spacecraft
cryogenics and weather predictions, whether to recycle for May 10 (T-O
hrs) or May 11 (I424 hrs). If launched on May 11, the flight plan will be
similar for the May 10 mission but the Sun elevation angle at lunar
landing will be 18.5 degrees instead of 7.8 degrees.
LAUNCH, MISSION TRAJECTORY AND MANEUVER DESCRIPTION
The information presented here is based on an on-time April 11 launch and
is subject to change before or during the mission to meet changing
conditions.
Launch
A Saturn V launch vehicle will lift the Apollo 13 spacecraft from Launch
Complex 39A, NASA-Kennedy Space Center, Fla.  The azimuth may vary from 72
to 96 degrees, depending on the time of launch.  The azimuth changes with
launch time to permit a fuel-optimum injection from Earth parking orbit to
a free-return circumlunar trajectory.
April 11 launch plans call for liftoff at 2:13 p.m. EST on an azimuth of
72 degrees. The vehicle will reach an altitude of 36 nautical miles
before first stage cutoff 51 nm downrange. During the 2 minutes 44 seconds
of powered flight, the first stage will increase vehicle velocity to 7,775
feet per second.* First stage thrust will reach a maximum of 8,995,108
pounds before center engine cutoff. After engine shutdown and separation
from the second stage, the booster will fall into the Atlantic Ocean about
364 nm downrange from the launch site (3O degrees North latitude and 74
degrees west longitude) about 9 minutes 4 seconds after liftoff.
The second stage (S-II) will carry the space vehicle to an altitude of 102
nm and a distance of 892 nm downrange. At engine shutdown, the vehicle
will be moving at a velocity of 21,508 fps. The four outer J-2 engines
will burn 6 minutes 32 seconds during the powered phase, but the center
engine will be cut off 4 minutes 47 seconds after S-II ignition.
At outboard engine cutoff, the S-II will separate and, following a
ballistic trajectory, plunge into the Atlantic about 2,450 nm downrange
from the Kennedy Space Center (31 degrees North latitude and 33.4 degrees
West longitude) some 20 minutes 41 seconds after liftoff.
The single engine of the Saturn V third stage (S-IVB) will ignite about 3
seconds after the S-II stage separates. The engine will fire for 143
seconds to insert the space vehicle into a circular Earth parking orbit of
103 nm beginning about 1,468 nm downrange.  Velocity at Earth orbital
insertion will be 24,243 fps at 11 minutes 55 second ground elapsed time
(GET). Inclination will be 33 degrees to the equator.
*NOTE: Multiply nautical miles by 1.1508 to obtain statute miles, multiply
feet per second by 0.6818 to obtain statute miles per hour.
The crew will have a backup to launch vehicle guidance during
powered-flight. If the Saturn instrument unit inertial platform fails, the
crew can switch guidance to the command module systems for first-stage
powered flight automatic control. Second and third stage backup guidance
is through manual takeover in which spacecraft commander hand controller
inputs are fed through the command module computer to the Saturn
instrument unit.
Earth Parking Orbit (EPO)
Apollo 13 will remain in Earth parking orbit for one and one-half
revolutions. The final "go" for the TLI burn will be given to the crew
through the Carnarvon, Australia, Manned Space Flight Network station.
Translunar Infection (TLI)
Midway through the second revolution in Earth parking orbit, the S-IVB
third-stage engine will restart at 2:35 GET over the mid-Pacific Ocean
near the equator and burn for almost six minutes to inject Apollo 13
toward the Moon. The velocity will increase from 25,593 fps to 36,030 fps
at TLI cutoff to free return circumlunar trajectory from which midcourse
corrections could be made with the SM RCS thrusters.
Transposition, Docking, and Ejection (TD&E)
After the TLI burn, the Apollo 13 crew-will separate the command/service
module from the spacecraft module adapter (SLA), thrust out away from the
S-VB, turn around and move back in for docking with the lunar module.
Docking should take place at about three hours and 21 minutes GET. After
the crew confirms all docking latches solidly engaged, they will
connect the CSM-to-LM umbilicals and pressurize the LM with oxygen from
the command module surge tank. At about 4:00 GET, the spacecraft will be
ejected from the spacecraft LM adapter by spring devices at the four LM
landing gear "knee" attach points. The ejection springs will impart about
one fps velocity to the spacecraft. A 9.4 fps S-IVB attitude thruster
evasive maneuver in plane at 4:19 GET will separate the spacecraft to a
safe distance from the S-IVB.
Saturn Third Stage Lunar Impact
Through a series of pre-set and ground-commanded operations, the S-IVB
stage/instrument unit will be-directed to hit the Moon within a target
area 375 nautical miles in diameter, centered just east of Lasberg D
Crater (3 degrees South latitude; 30 degrees West longitude),
approximately 124 miles west of the Apollo 12 landing site.
The planned impact will provide a seismic event for the passive
seismometer experiment placed on the lunar surface by the Apollo 12
astronauts in November 1969.
The residual propellants in the S-IVB will be used to attempt the lunar
impact. Part of the remaining liquid oxygen (LOX) will be dumped through
the engine for 48 seconds to slow the vehicle into a lunar impact
trajectory. The liquid hydrogen tank's continuous venting system will vent
for five minutes.
A mid-course correction will be made with the stage's auxiliary propulsion
system (APS) ullage motors. A second APS burn will be used if necessary,
at about 9 hours GET, to further adjust the impact point. Burn time and
attitude will be determined from onboard systems and tracking data
provided to ground controllers by the Manned Space Flight Network.
The LOX dump by itself would provide a lunar impact, the mid-
course correction burns will place the S-IVB/IU within the desired
target area for impact about 20 minutes after the command/service
module enters lunar orbit.
The schedule of events concerning the lunar impact is:
  Time
Hrs: Min				Event
 02  42 			Translunar injection (TLI) --
  				maneuver completion
 04  19 			Begin S-IVB evasive maneuver (APS
  				engines)
 04  21 			End evasive maneuver
 04  36 			LH2 tank continuous vent on
 04  41 			Begin LOX dump
 04  41 			LH2 tank continuous vent off
 04  42 			End L0X dump
 06  00 			Begin first APS burn
 08  59 			Begin final APS burn (if required) -
 09  04 			APS ullage engines off
 77  46 			Lunar impact of S-IVB/IU
Translunar Coast
Up to four midcourse correction burns are planned during the spacecraft's
translunar coast, depending upon the accuracy of the trajectory resulting
from the TLI maneuver. If required, the midcourse correction burns are
planned at TLI+9 hours, TLI+ 30 hours, 41 minutes, lunar orbit insertion
(LOI)-22 hours and LOI-5 hours. The MCC-2 is a 15 fps SPS hybrid transfer
maneuver which lowers pericynthion from 210 nm to 59 nm and places Apollo
13 on a non-free-return trajectory.
Return to the free-return trajectory is always within the capability of
the spacecraft service propulsion or descent propulsion systems.
During coast periods between midcourse corrections, the spacecraft will be
in the passive thermal control (PTC) or "barbecue" mode in which the
spacecraft will rotate slowly about its roll axis to stabilize spacecraft
thermal response to the continuous solar exposure.
Lunar Orbit Insertion (LOI)
The lunar orbit insertion burn will be made at 77:25 GET at an altitude of
about 85 nm above the Moon. The LOI burn will have a nominal retrograde
velocity change of 2815 fps and will insert Apollo 13 into a 57x168 nm
elliptical lunar orbit.
Descent Orbit Insertion (DOI)
A 213 fps SPS retrograde burn at 81:45 GET will place the CSM LM into a
757 nm lunar orbit from which the LM will begin the later powered descent
to landing. In Apollos 11 and 12, DOI was a separate maneuver using the LM
descent engine. The Apollo 13 DOI maneuver in effect is a combination
LOI-2 and DOI and produces two benefits: conserves LM descent propellant
that would have been used or DOI and makes this propellant available for
additional hover time near the surface, and allows 11 lunar revolutions of
spacecraft tracing in the descent orbit to enhance position/velocity
(state vector) data for updating the LM guidance computer during the
descent and landing phase.
Lunar Module Separation
The lunar module will be manned and checked out or undocking and
subsequent landing on the lunar surface north of the crater, Fra Mauro.
Undocking during the 12th revolution will take place at 99:16 GET. A
radically downward service module RCS burn of 1 fps will place the CSM on
an equiperiod orbit with a maximum separation of 2.5 nm.
CSM Circularization:
During the 12th revolution, a 70 fps posigrade SPS burn at 100:35 GET will
place the CSM into 52x62 nm lunar orbit, which because of perturbations of
the lunar gravitational potential, should become nearly circular at the
time of rendezvous with the LM.
Power Descent Initiation (PDI), Lunar Landing
During the 14th revolution a three-phase powered descent (PD) maneuver
begins at pericynthion at 103:31 GET using the LM descent-engine to brake
the vehicle out of the descent orbit.  The guidance-controlled PD maneuver
starts about 260 nm prior to touchdown, and is in retrograde attitude
to reduce velocity to essentially zero at the time vertical descent
begins. Spacecraft attitude will be windows up from powered descent
initiation to the end of the braking phase so that the LM landing radar
data can be integrated continually by the LM guidance computer and better
communications can be maintained. The braking phase ends at about 7,400
feet above the surface and the spacecraft is rotated more toward an
upright windows-forward attitude to permit a view of the landing site. The
start of the approach phase is called high gate, and the start of the
landing phase at about 500 feet is called low gate.
Both the approach (visibility) phase and landing phase allow pilot
takeover from guidance control as well as visual evaluation of the landing
site. The final vertical descent to touchdown begins at about 100 feet
when all forward velocity is nulled out. Vertical descent rate will be 3
fps. The crew may elect to take over manual control at approximately 500
feet. The crew will be able to return to automatic landing control after a
period of manned maneuvering if desirable.  Touchdown will take place at
103:42 GET.
Lunar Surface Exploration
During the 33 1/2 hours Apollo 13 commander James Lovell and lunar module
pilot Fred Haise are on the surface, they will leave the lunar module
twice for four-hour EVAs. These are extendible to five hours in real time
if the physical conditions of the astronauts and amount of remaining
consumables permit.
In addition to gathering more data on the lunar environment and bringing
back geological samples from a third lunar landing site, Lovell and Haise
will deploy a series of experiments which will relay back to Earth
long-term scientific measurements of the Moon's physical and environmental
properties.
The experiments series, called the Apollo Lunar Surface Experiment Package
(ALSEP), will be left on the surface and could transmit scientific and
engineering data to the Manned Space Flight Network for at least a year.
The ALSEP for Apollo 13, stowed in the LM descent stage scientific
equipment bay, comprises components for the five ALSEP experiments --
passive seismic, heat flow, charged particle lunar environment, cold
cathode gauge, and lunar dust detector. These experiments are aimed
toward determining the structure and state of the lunar interior, the
composition and structure of the lunar surface and processes which modify
the surface, and evolutionary sequence leading to the Moon's present
characteristics. The Passive Seismic Experiment will become the second
point in a lunar seismic net begun with the first ALSEP at the Surveyor
III landing site of Apollo 12. Those two seismometers must continue to
operate until the next seismometer is emplaced to complete the set. The
heat flow experiment includes drilling two 10-foot holes with the lunar
surface drill.
While on the surface, the crew's operating radius will be limited by the
range provided by the oxygen purge system (OPS), the reserve backup for
each man's portable life support system (PLSS) backpack.  The OPS supplies
45 minutes of emergency breathing oxygen and suit pressure.
Among other tasks assigned to Lovell and Haise for the two EVA
periods are:
*  Collect a contingency sample of about two pounds of lunar material.
*  Gather about 95 pounds of representative lunar surface material,
including core samples, individual rock samples and fine-grained fragments
from the Fra Mauro hilly uplands site. The crew will photograph thoroughly
the areas from which samples are taken.
*  Make observation's and gather data on the mechanical properties and
terrain characteristics of the lunar surface and conducting other lunar
field geological surveys, including digging a two-foot deep trench for a
soil mechanics investigation.
*  Photograph with the lunar stereo closeup camera small geological features
that would be destroyed in any attempts to gather them for return to
Earth.
*  Deploy and retrieve a windowshade-like solar wind-composition experiment
similar to the ones used in Apollos 11 and 12.
Early in the first EVA, Lovell and Haise will set up the erectable S-Band
antenna near the LM for relaying voice, TV, and LM telemetry to MSFN
stations. After the antenna is deployed, Haise will climb back into the LM
to switch from the LM steerable S-Band antenna to the erectable antenna
while Lovell makes final adjustments to the antenna's alignment. Haise
will then rejoin Lovell on the lunar surface to set up a United States
flag and continue with EVA tasks. Red stripes around the elbows and knees
of Lovell's pressure suit will permit crew recognition during EVA
television transmissions and on photographs.
Ascent, Lunar Orbit Rendezvous
Following the 33-hour lunar stay the LM ascent stage will lift off the
lunar surface to begin the rendezvous sequence with the orbiting CSM.
Ignition of the LM ascent engine will be at 137:09 for a seven minute
eight second burn attaining a total velocity of 6,044 fps.  Powered ascent
is in two phases: vertical ascent for terrain clearance and the orbital
insertion phase. Pitchover along the desired launch azimuth begins as the
vertical ascent rate reaches 50 fps about 10 seconds after liftoff at
about 272 feet in altitude. Insertion into a 9x44 nm lunar orbit will take
place about 166 nm west of the landing site.
Following LM insertion into lunar orbit, the LM crew will compute onboard
the major maneuvers for rendezvous with the CSM which is about 267 nm
ahead of and 51 miles above the LM at this point. All maneuvers in the
sequences will be made with the LM RCS thrusters.  The pre-mission
rendezvous sequence maneuvers, time, and velocities, which likely will
differ slightly in real time, are as follows:
Concentric sequence initiate (CSI): At first LM apolune after insertion,
138:19 GET, 50 fps posigrade, following some 20 minutes of LM rendezvous
radar tracking and CSM sextant/VHF ranging navigation.  CSI will be
targeted to place the LM in an orbit 15 nm below the CSM at the time of
the later constant delta height (CDH) maneuver (139:04).
The CSI burn may also initiate corrections for any out-of plane
dispersions resulting from insertion azimuth errors. The resulting LM
orbit after CSI will be 45x43.5 nm and will have a catch-up rate to the
CSM of about 120 feet per second.
Terminal phase initiation (TPI): This maneuver occurs at 139:46 and adds
24.7 fps along the line of sight toward the CSM when the elevation angle
to the CSM reaches 26.6 degrees. The LM orbit becomes 61x44 nm and the
catch-up rate to the CSM decreases to a closing rate of 133 fps.
Midcourse correction maneuvers will be made if needed, followed by four
braking maneuvers. Docking nominally will take place at 140:25 GET to end
the three and one-half hour rendezvous sequence.
The LM ascent stage will be jettisoned at 143:04 GET and a CSM RCS 1.0 fps
maneuver will provide separation. Prior to transferring to the command
module, the LM crew will set up the LM guidance system to maintain the
ascent stage in an inertial attitude. At about 144:32 GET the LM RCS
thrusters will ignite on ground command for 186 fps retrograde burn
targeted for ascent stage impact at 145:00 about 35 miles from the landing
site. The burn will have a small out-of-plane north component so that the
ground track will include the original landing site. The ascent stage will
impact at about 5508 fps at an angle of four degrees relative to the local
horizontal. The ascent stage deorbit serves to remove debris from lunar
orbit. Impacting an object with a known velocity and mass near the landing
site will provide experimenters with an event for calibrating readouts
from the ALSEP seismometer left behind.
A plane change maneuver at 154:13 GET will place the CSM on an orbital
track passing directly over the crater Descartes and Davy Rille eight
revolutions later for photographs from orbit. The maneuver will be a 825
fps/SPS burn out of plane for a plane change of 8.8 degrees, and will
result in an orbit inclination of 11.4 degrees.
Transearth Injection (TEI)
The nominal transearth injection burn will be at 167:29
GET following 90 hours in lunar orbit. TEI will take place on
the lunar farside, will be a 3,147 fps posigrade SPS burn of
two minutes 15 seconds duration and will produce an entry
velocity of 36,129 fps after a 72 hours transearth flight time.
Transearth Coast
Three entry corridor-control transearth midcourse correction burns will
be as if needed: MCC-5 at TEI+15 hours, MCC-6 at entry interface (EI) 22
hours and NCC-7 at EI-3 hrs.
Entry, Landing
Apollo 13 will encounter the Earth's atmosphere (400,000 feet) at 240:50
GET at a velocity of 36,129 fps and will land approximately 1,250 nm
downrange from the entry-interface point using the spacecraft's lifting
characteristics to reach the landing point.  Splashdown will be at 241:04
at 1.5 degrees South latitude by 157.5 degrees West longitude.
Recovery Operations
Launch abort landing areas extend downrange 3,400 nautical miles from
Kennedy pace Center, Panwise 50 nm miles above and below the limits of the
variable launch azimuth (72-96 degrees) in the Atlantic Ocean. On station
in the launch abort area will be the destroyer USS New.
The landing platform-helicopter (LPH) Iwo Jima, Apollo 13 prime recovery
ship, will be stationed near the Pacific Ocean end-of-mission aiming
point prior entry.
Splashdown for a full-duration lunar landing mission launched on time
April 11 will be at one degree 34 minutes South by 17 degrees 30 minutes
West about 180 nautical miles South of Christmas Island, at 241:04 GET
(3:17 p.m. EST) April 21.
In addition to the primary recovery vessel located on the mid-Pacific
recovery line and the surface vessel in the launch abort area, eight
HC-130 aircraft will be on standby at five staging bases around the Earth:
Guam; Hawaii; Azores; Ascension Island; and Florida.
Apollo 13 recovery operations will be directed from the Recovery
Operations Control Room in the Mission Control Center, supported by the
Atlantic Recovery Control Center, Norfolk, Va., and the Pacific Recovery
Control Center, Kunia, Hawaii.
After splashdown, the Apollo 13 crew will don clean coveralls and filter
masks passed to them through the spacecraft hatch by a recovery swimmer.
The crew will be carried by helicopter to the Iwo Jima where they will
enter a Mobile Quarantine Facility (MQF) about 90 minutes after landing.
APOLLO 13 ONBOARD TELEVISION
Apollo 13 will carry two color and one black-and-white television cameras.
One color camera will be used for command module cabin interiors and
out-the-window Earth/Moon telecasts, and the other color camera will be
stowed in the LM descent stage from where it will view the astronaut
initiate egress to the lunar surface and later will be deployed on a
tripod to transmit a real-time picture of the two periods of lunar surface
EVA. The black-and-white camera will be carried in the LM cabin. It will
only be used as a backup to the lunar surface color camera.
The two color TV cameras are essentially identical, except for additional
thermal protection on the lunar surface camera. Built by Westinghouse
Electric Corp., Aerospace Division, Baltimore, Md., the color cameras
output a standard 525-line, 30 frame-per-second signal in color by use of
a rotating color wheel system.
The color TV cameras weigh 12 pounds and are fitted with zoom lenses for
wide-angle or closeup fields of view. The CM camera is fitted with a
three-inch monitor for framing and focusing. The lunar surface color camera
has 100 feet of cable available.
The backup black-and-white lunar surface TV camera, also built by
Westinghouse, is of the same type used in the first manned lunar landing
in Apollo 11. It weighs 7.25 pounds and draws 6.5 watts of 24-32 volts DC
power. Scan rate is 10 frames-per second at 325 lines-per-frame. The
camera body is 10.6 inches long, 6.5 inches wide and 3' inches deep, and
is fitted with bayonet-mount wide-angle and lunar day lenses.
During the two lunar surface EVA periods, Apollo 13 commander Lovell will
be recognizable by red stripes around the elbows and knees of his pressure
suit.
The following is a preliminary plan for TV transmissions based upon a
2:13 p.m. EST April 11 launch.
APOLLO 13 SCIENCE
Lunar Orbital Photography
Science experiments and photographic tasks will be conducted from the CSM
during the Apollo 13 mission. During the translunar phase of the mission,
photography will be taken of the Earth as well as various operational
photography.
During lunar orbit, various lunar surface features including candidate
landing sites Censorinus, Descartes and Davy Rille and the Apollo 11 and
12 landing sites will be photographed with the Lunar Topographic Camera.
In addition, five astronomical phenomena will be photographed:
1) Photographs will be taken of the solar corona using the Moon as an
occulting edge to block out the solar disk.
2) Photography will be taken of the zodiacal light which is believed to
originate from reflected sunlight in the asteroid belt. Earth observation
of zodiacal light is inconclusive due to atmospheric distortion.
3) Photography will be taken of lunar limb brightening, which appears as
bright rim light above the horizon following Lunar sunset.
4) Photographs will be taken of the Comet J.C. Bennett, 1969i which should
be visible from lunar orbit during the Apollo 13 mission.
5) Photographs will be taken of the region of Gegenschein which is a faint
light source covering a 20 degree field of view about the Earth-Sun line
on the opposite side of the Earth from the Sun (anti-solar axis).  One of
the theories for the Gegenschein source is the existence of trapped
particles of matter at the Moulton point which produce brightness due to
reflected sunlight. The Moulton point is a theoretical point located
94O,000 statute miles from the Earth along the antisolar axis at which the
sum of all gravitational forces is zero.  From the vantage point of lunar
orbit, the Moulton point region may be photographed from approximately 15
degrees off the Earth/Sun line.  These photographs should show if
Gegenschein results from the Moulton point theory or from zodiacal light
or a similar source.
Photographic studies will be made on Apollo 13 of the ice particle flow
following a water dump and of the gaseous cloud which surrounds a manned
spacecraft in a vacuum and results from liquid dumps, outgassing, etc.
In addition to the photographic studies, an experiment will be conducted
with the CSM VHF communications link. During this experiment, the VHF
signal will be reflected from the lunar surface and received by a 150-foot
antenna on Earth. By analysis of the wavelength of the received signal,
certain lunar subsurface characteristics may be discernible such as the
depth of the lunar regolith layer. This experiment is called VHF Bistatic
Radar.
Charged Particle Lunar Environment Experiment (CPLEE)
The scientific objective of the Charged Particle Lunar Environment
Experiment is to measure the particle energies of protons and electrons
that reach the lunar surface from the Sun. Increased knowledge on the
energy distribution of these particles will help us understand how they
perturb the Earth/Moon system. At some point electrons and protons in the
magnetospheric tail of the Earth are accelerated and plunge into the
terrestrial atmosphere causing the spectacular auroras and the Van Allen
radiation. When the Moon is in interplanetary space the CPLEE measures
proton and electrons from solar fairies which results in magnetic storms
in the Earth's atmosphere. Similar instruments have been flown on Javelin
rockets and on satellites. The lunar surface, however, allows data to be
gathered over a long period of time and from a relatively stable platform
in space.
To study these phenomena, the CPLEE measures the energy of protons and
electrons simultaneously from 50 electron volts to 50,000 electron volts
(50Kev)- The solar radiation phenomena measured are as follows:
a. Solar wind electrons and protons 50ev-5Kev.
b. Thermalized solar wind protons and electrons 5Oev-lOK ev.
c. Magnetospheric tail particles 50ev to 50Kev.
d. Low energy solar cosmic rays 40ev-50Kev.
This experiment is distinct from the ALSEP Solar Wind Spectrometer (SWS)
flown on Apollo 12 which measures direction as well as energy levels. The
SWS measures electrons from 1 or 5ev to 1,400ev and protons from 75ev to
lO,OOOev.
The detector package contains two spectrometers providing data on the
direction of the incoming flux.
Each spectrometer has six particle detectors: five C-shaped channeltron
photon-multipliers and one funneltron, a helical shaped photon multiplier.
Particles of a given charge and different energies on entering the
spectrometer are subject to varying voltages and deflected toward the five
channeltrons while particles of the opposite charge are deflected toward
the funneltron. Thus electrons and protons are measured simultaneously in
six different energy levels. The voltages are changed over six steps; + or
- 35Y, + or - 350 volts and + or - 500V. In this way electrons and protons
are measured from 50ev to 70Kev in a period of less than 20 seconds.
The channeltron is a glass capillary tube having an inside diameter of
about one millimeter and a length of 10 centimeters.  The helical
funneltron has an opening of 8mm. When a voltage is applied between the
ends of the tube, an electric field is established down its length.
Charged particles entering the tube are amplified by a factor of lO^8.
The spectrometers have two ranges of sensitivity and can measure fluxes
between 10^4 and 10^10 particles/cm2-sec-steradian.
The charged particle lunar environment experiment (CPLEE) and data
analysis are the responsibility of Dr. Brian O'Brien, University of Sydney
(Australia) and Dr. David Reasoner, Rice University, with Dr. O'Brien
assuming the role of Principal Investigator.
Lunar Atmosphere Detector (LAD)
Although the Moon is commonly described as a planetary body with no
atmosphere, the existence of some atmosphere cannot be doubted.  Two
sources of this atmosphere are predicted: internal, i.e., degassing from
the interior of the Moon either by constant diffusion through its surface
or intermittent release from active vents; external i.e., solar wind and
vaporization during meteorite impacts. Telescopic observations from
polarized scattered light indicate that the atmospheric pressure could not
exceed one millionth of a torr (a torr is defined as 1/760 of the standard
atmosphere).
Measurements will be of the greatest significance if it turns out through
later orbital sensors that they are of internal origin. The Earth's
atmosphere and oceans have been released from the Earth's interior by
degassing. The most certain source, however, is the solar wind whose
ionized particles become neutralized in the lunar atmosphere and then are
released as neutral gases. Neon is the predominant gas expected. Lighter
gases such as hydrogen and helium escape and heavier ones statistically
should be present in small quantities. Neutral particles are ionized in he
lunar atmosphere, further reducing the numbers present; others will escape
as the temperature rises (and concentrate near the surface when it falls).
The LAD utilizes a cold cathode ionization gauge to measure the density of
neutral particles at the lunar surface and the variations in density
association with lunar phase or solar activity.  The ionization gauge is
basically a crossed electro-magnetic field device. Electrons in the gauge
are accelerated by the combined magnetic and electric fields producing a
collision are collected by the cathode where they form a flow of positive
ions. The positive ions current is found to be proportional to the density
of the gas molecules entering the gauge. In addition, the gauge
temperature is read over the range of -90 degrees to 120 degrees C with +
or - 5 degrees C accuracy.
From the density and temperature data the pressure of the ambient lunar
atmosphere can then be calculated. Chemical composition of the atmosphere
however is not directly measured but the gauge has been calibrated for
each gas it is expected to encounter on the lunar surface and some
estimates can be made of the chemical composition.  Any one of seven
different dynamic ranges may be selected permitting detection of neutral
particles from 10^-6 Torr (highest pressure predicted) to 10^-12 Torr
(maximum capability of gauge).  For pressure greater than 10^lO Torr
accuracies of + or - 30% will be obtained, for pressures less then 10^-1O
Torr accuracies + or - 50% will be obtained. The experiment, therefore,
will reduce the present uncertainty from a magnitude o a factor.
The Lunar Atmosphere Detector (LAD) and data are the responsibility of
Francis Johnson, University of Texas (Dallas) and Dallas Evans, Manned
Spacecraft Center, with Dr. Johnson serving as Principal Investigator.
Lunar Heat Flow Experiment (HFE)
The scientific objective of the Heat Flow experiment is to measure the
steady-state heat flow from the Lunar interior. Two predicted sources of
heat are: l) original - heat at the time of the Moon's formation-and 2)
radioactivity. Scientists believe that heat could have been generated by
the infalling of material and its subsequent compaction as the Moon was
formed. Moreover, varying amounts of the radioactive elements uranium,
thorium and potassium were found present in the Apollo 11 and 12 lunar
samples which if present at depth, would supply significant amounts of
heat. No simple way has been devised for relating the contribution of each
of these sources to the present rate of heat loss. In addition to
temperature, the experiment is capable of measuring the thermal
conductivity of the lunar rock material.
The combined measurement of temperature and thermal conductivity gives
the net heat flux from the lunar interior through the lunar surface.
Similar measurements on Earth have contributed basic information to our
understanding of volcanoes, earthquakes and mountain
building processes. In conjunction with the seismic and magnetic data
obtained on other lunar experiments the values derived from the heat flow
measurements will help scientists to build more exact models of the Moon
and thereby give us a better understanding of its origin and history. -
The Heat Flow experiment consists of instrument probes, electronics and
emplacement tool and the lunar surface drill. Each of two probes is
connected by a cable to an electronics box which rests on the lunar
surface. The electronics, which provide control, monitoring and data
processing for the experiment, is connected to the ALSEP central station.
Each probe consists of two identical 20-inch (50 cm) long sections each of
which contains a "gradient" sensor bridge, a ring sensor bridge and two
heaters. Each bridge consists of four platinum resistors mounted in a
thin-walled fiberglass cylindrical shell. Adjacent areas of the bridge are
located in sensors at opposite ends of the 20-inch fiberglass probe
sheath. Gradient bridges consequently measure the temperature difference
between two sensor locations.
In thermal conductivity measurements at very low values a heater
surrounding the gradient sensor is energized with 0.002 watts and the
gradient sensor values monitored. The rise in temperature of the gradient
sensor is a function of the thermal conductivity of the surrounding lunar
material. For higher range of values, the heater is energized at 0.5 watts
of heat and monitored by a ring sensor. The rate of temperature rise,
monitored by the ring sensor is a function of the thermal conductivity of
the surrounding lunar material. The ring sensor, approximately four inches
from the heater, is also a platinum resistor. A total of eight thermal
conductivity measurements can be made. The thermal conductivity mode of
the experiment will be implemented about twenty days (500 hours) after-
deployment. This is to allow sufficient time for the perturbing effects of
drilling and emplacing the probe in the borehole to decay;  i.e., for the
probe and casings to come to equilibrium with the lunar subsurface.
A 30-foot (10 meter) cable connects each probe to the electronics box. In
the upper six feet of the borehole the cable contains four evenly spaced
thermocouples: at the top of the probe; at 26" (65 cm), 45" (115 cm), and
66" (165 cm). The thermocouples will measure temperature transients
propagating downward from the lunar surface. The reference junction
temperature for each thermocouple is located in the electronics box. In
fact, the feasibility of making a heat flow measurement depends to a large
degree on the low thermal conductivity of the lunar surface layer, the
regolith. Measurement of lunar surface temperature variations by
Earth-based telescopes as well as the Surveyor and Apollo missions show a
remarkably rapid rate of cooling. The wide fluctuations in temperature of
the lunar surface (from -250 degrees F to +250 degrees) are expected to
influence only the upper six feet and not the bottom 3 feet of the
borehole.
The astronauts will use the Apollo Lunar Surface Drill (ALSD) to make a
lined borehole in the lunar surface for the probes. The drilling energy
will be provided by a battery powered rotary percussive power head. The
drill rod consists of fiberglass tubular sections reinforced with boron
filaments (each about 20 inches or 50 cm long). A closed drill bit, placed
on the first drill rod, is capable of penetrating the variety of rock
including three feet of vesicular basalt (40 per cent porosity). As lunar
surface penetration progresses, additional drill rod sections will be
connected to the drill string. The drill string is left in place to serve
as a hole casing.
An emplacement tool is used by the astronaut to insert the probe to full
depth. Alignment springs position the probe within the casing and assure a
well-defined radiative thermal coupling between the probe and the
borehole. Radiation shields on the hole prevent direct sunlight from
reaching the bottom of the hole.
The astronaut will drill a third hole near the HFE and obtain cores of
lunar material for subsequent analysis of thermal properties.
Heat flow experiment, design and data analysis are the responsibility of
Dr. Marcus Langseth of the Lamont-Doherty Geological Observatory; Dr.
Sydney Clark, Jr., Yale University, and Dr. M. G.  Simmons, MIT; with Dr.
Langseth assuming the role of Principal Investigator.
Passive Seismic Experiment (PSE)
The ALSEP Passive Seismic Experiment (PSE) will measure seismic activity
of the Moon and obtain information on the physical properties of the lunar
crust and interior. The PSE will detect surface tilt produced by tidal
deformations, moonquakes and meteorite impacts.
The passive seismometer design and subsequent experiment analysis are the
responsibility of Dr. Gary Latham of the Lamont-Doherty Geological
Observatory.
A similar passive seismic experiment was deployed as part of the Apollo 12
ALSEP station at Surveyor crater last November and has transmitted
Earthward lunar surface seismic activities since that time. The Apollo 12
and 13 seismometers differ from the seismometer left at Tranquillity Base
in July 1969 by the Apollo 11 crew in that they are continuously powered
by a SNAP-27 radioisotope electric generator, while the Apollo 11
seismometer was powered by solar energy and could output data only during
the lunar day at its location.
After Lovell and Haise ascend from the lunar surface and rendezvous with
the command module in lunar orbit, the lunar module ascent stage will be
jettisoned and later ground commanded to impact on the lunar surface about
42 statute miles from the Apollo 13 landing site at Fra Mauro. Impact of
an object of known mass and velocity will assist in calibrating the Apollo
13 seismometer readouts as well as providing comparative readings between
the Apollo 12 and 13 seismometers forming the first two stations of a
lunar surface seismic network.
There are three major physical components of the PSE:
*  The sensor assembly consists of three, long-period seismometers with
orthogonally-oriented, capacitance type seismic sensors, measuring along
two horizontal axes and one vertical axis. This is mounted on a gimbal
platform assembly. There is one short period seismometer which has
magnet-type sensors. It is located directly on the base of the sensor
assembly.
*  The leveling stool allows manual leveling of the sensor assembly by the
astronaut to within + or - 5 degrees, and final leveling to within 3 arc
seconds by control. motors.
*  The thermal shroud covers and helps stabilize the temperature of the
sensor assembly. Also, two radioisotope heaters will protect the
instrument from the extreme cold of the lunar night.
Solar Wind Composition Experiment (SWCE)
The scientific objective of the solar wind composition experiment is to
determine the elemental and isotopic composition of the noble gases in the
solar wind.  The solar wind composition detector experiment design and
subsequent data analysis are the responsibility of J. Geiss and P.
Eberhardt, University of Bern (Switzerland) and P. Signer, Swiss Federal
Institute of Technology, with Professor Geiss assuming the responsibility
of Principal Investigator.
As in Apollo 11 and 12 the SWC detector will be deployed on the Moon and
brought back to Earth by the astronauts. The detector, however, will be
exposed to the solar wind flux for 20 hours instead of two hours as in
Apollo 11 and 18 hours 42 minutes on Apollo 12.
The solar wind composition detector consists of an aluminum foil four
square feet in area and about 0.5 mils thick rimmed by Teflon for
resistance to tear during deployment. A staff and yard arrangement will be
used to deploy the foil and to maintain the foil approximately
perpendicular to the solar wind flux. Solar wind particles will penetrate
into the foil while cosmic rays will pass right through. The solar wind
particles will be firmly trapped at a depth of several hundred atomic
layers. After exposure on the lunar surface, the foil is reeled and
returned to Earth.
Dust Detector
The ALSEP Dust Detector is an engineering measurement designed to detect
the presence of dust or debris that may impinge on the ALSEP or accumulate
during its operating life.
The measurement apparatus consists of three calibrated solar cells, one
pointing in east, west and vertical to face the elliptic path of the
Sun. The detector is located on the central station.
Dust accumulation on the surface of the three solar cell will reduce the
solar illumination detected by the cells. The temperature of each cell
will be measured and compared with predicted values.
Field Geology Investigations
The scientific objectives of the Apollo Field Geology Investigations are
to determine the composition of the Moon and the processes which shape its
surfaces. This information will help to determine the history of the Moon
and its relationship to the Earth. Apollo 11 visited the Sea of
Tranquillity (Mare Tranquillitatis) and Apollo 12 studied the Ocean of
Storms (Oceanus Procellarum). The results of these studies should help
establish the nature of Mare-type areas.  Apollo 13 will investigate a
hilly upland area.
Geology investigation of the Moon actually began with the telescope.
Systematic geology mapping began 10 years ago with a team of
scientists at the U.S. Geological Survey. Ranger, Surveyor, and
especially Lunar Orbiter data enormously increased the detail and
accuracy of these studies. The Apollo 11 and 12 investigations
represent another enormous advancement in providing new
evidence on the Moon's great age, its curious chemistry, the
surprisingly high density of the lunar surface material.
On Apollo 13, almost the entire second EVA will be devoted to the
Field Geology Investigations and the collection of documental
samples. The sample locations will be carefully photographed before
and after sampling. The astronauts will carefully describe the setting
from which the sample is collected. In addition to specific tasks, the
astronauts will be free to photograph and sample phenomena which
they judge to be unusual, significant, and interesting. The astronauts
are provided with a package of detailed photo maps which they will
use for planning traverses. Photographs will be taken from the LM
window. Each feature or family of features will be described, relating
to features on the photo maps. Areas and features where
photographs should be taken and representative samples collected
will be marked on the maps. The crew and their ground support
personnel will consider real-time deviation from the nominal plan
based upon an on-the-spot analysis of the actual situation. A trench
will be dug for soil mechanics investigations.
The Earth-based geologists will be available to advise the astronauts
in real-time and will work with the data returned, the photos,
the samples of rock and the astronauts' observations to reconstruct
here on Earth the astronauts traverse on the Moon.
Each astronaut will carry a Lunar Surface Camera (a modified 70 mm
electric Hasselblad). The camera has a 60 mm lens and a Reseau
plate. Lens apertures range from 5/5.6 to f/45. Its focus range is
from three feet to infinity. A removable polarizing filter is attached
to the lens of one of the cameras and can be rotated in 45-degree
increments for light polarizing studies.
A gnomon, used for metric control of near field (less than 10 feet)
stereoscopic photography, will provide angular orientation relative to
the local vertical. Information on the distances-to objects and on the
pitch, roll, and azimuth of the camera's optic axis are thereby
included in each photograph. The gnomon is a weighted tube
suspended vertically on a tripod supported gimbal. The tube extends
one foot above the gimbal and is painted with a gray scale in bands
one centimeter wide. Photogrammetric techniques will be used to
produce three-dimensional models and maps of the lunar surface
from the angular and distance relationship between specific objects
recorded on the film.
The 16 mm Data Acquisition Camera will provide times sequence
coverage from within the LM. It can be operated in several automatic
modes, ranging from one frame/second to 24 frames/second. Shutter
speeds, which are independent of the frame rates, range from
1/1000 second to 1/60 second. Time exposures are also possible.
While a variety of lenses is provided, the 18 mm lens will be used to
record most of the geological activities in the one frame/second
mode. A similar battery powered 16 mm camera will be carried in
EVA.
The Lunar Surface Close-up Camera will be used to obtain very high
resolution close-up stereoscopic photographs of the lunar surface to
provide fine scale information on lunar soil and rock textures. Up to
100 stereo pairs can be exposed on the preloaded roll of 35 mm color
film. The handle grip enables the astronaut to operate the camera
from a standing position. The film drive and electronic flash are
battery-operated. The camera photographs a 3"x3" area of the lunar
surface.
Geological sampling equipment includes tongs, scoop, hammer, and
core tubes. A 24-inch extension handle is provided for several of
the tools to aid the astronaut in using them without kneeling.
Sample return containers (SRC) have been provided for return of up
to 40 pounds each of lunar material for Earth-based analysis. The
SRC's are identical to the ones used on the Apollo 11 and 12 missions.
They are machined from aluminum forgings and are designed to
maintain an internal vacuum during the outbound and return flights.
The SRC's will be filled with representative samples of lunar surface
material, collected and separately bagged by the astronauts on their
traverse and documented by verbal descriptions and photography.
Subsurface samples will be obtained by using drive tubes 16 inches
long and one inch in diameter. A few grams of material will be
preserved under lunar vacuum conditions in a special environmental
sample container.
This container will be opened for analysis under vacuum conditions
equivalent to that at the lunar surface. Special containers are
provided for a magnetic sample and a gas analysis sample.
SNAP-27
SNAP-27 is one of a series of radioisotope thermoelectric generators,
or atomic batteries, developed by the U.S. Atomic Energy Commission
under its SNAP program. The SNAP (Systems for Nuclear Auxiliary
Power) Program is directed at development of generators and
reactors for use in space, on land, and in the sea.
SNAP-27 was first used in the Apollo 12 mission to provide
electricity for the first Apollo Lunar Surface Experiments Package
(ALSEP). A duplicate of the Apollo 12 SNAP-27 will power the Apollo
13 ALSEP.
The basic SNAP-27 unit is designed to produce at least 63 electrical watts
of power. It is a cylindrical generator fueled with the radioisotope
plutonium 238. It is about 18 inches high and 16 inches in diameter,
including the heat radiating fins. The generator, making maximum use of
the lightweight material beryllium, weighs about 28 pounds unfueled.
The fuel capsule, made of a superalloy material, is 16.5 inches long
and 2.5 inches in diameter. It weighs about 15.5 pounds, of which
8.36 pounds represent fuel. The plutonium 238 fuel is fully oxidized
and is chemically and biologically inert.
The rugged fuel capsule is contained within a graphite fuel cask from
launch through lunar landing. The cask is designed to provide
reentry heating protection and added containment for the fuel
capsule in the unlikely event of an aborted mission. The cylindrical
cask with hemispherical ends includes a primary graphite heat
shield, a secondary beryllium thermal shield, and a fuel capsule
support structure made of titanium and Inconel materials. The cask
is 23 inches long and eight inches in diameter and weighs about 24.5
pounds. With the fuel capsule installed, it weighs about 40 pounds. It
is mounted on the lunar module descent stage by a titanium support
structure.
Once the lunar module is on the Moon, the lunar module pilot will remove
the fuel capsule from the cask and insert it into the SNAP- 27 generator
which will have been placed on the lunar surface near the module.
The spontaneous radioactive decay of the plutonium 238 within the
fuel capsule generates heat in the generator. An assembly of 442
lead telluride thermoelectric elements converts this heat -- 1480
thermal watts -- directly into electrical energy -- at least 63 watts.
There are no moving parts.
Plutonium 238 is an excellent isotope for use in space nuclear generator.
At the end of almost 90 years, plutonium 238 will still supply half of its
original heat. In the decay process, plutonium 238 emits mainly the nuclei
of helium (alpha radiation)j a very mild type of radiation with a short
emission range.
Before the use of the SNAP27 system in the Apollo program was authorized a
thorough review was conducted to assure the health and safety of personnel
involved in the mission and the general public. Extensive safety analyses
and tests were conducted which demonstrated that the fuel would be safely
contained under almost all credible accident conditions.
Contractors for SNAP-27
General Electric Co., Missile and Space Division, Philadelphia, Pa.,
designed, developed, and fabricated the SNAP-27 generator for the
ALSEP.
The 3M Co., St. Paul, Minn., fabricated the thermoelectric elements
and assembled the SNAP-27 generator.
Solar Division of International Harvester, San Diego, Calif., fabricated
the generator's beryllium structure.
Hitco, Gardena, Calif., fabricated the graphite structure for the
SNAP-27 Graphite LM Fuel Cask.
Sandia Corp., a subsidiary of Western Electric, operator of AEC's
Sandal Laboratory, Albuquerque N.M., provided technical direction
for the SNAP-27 program.
Savannah River Laboratory, Aiken, S.C., operated by the DuPont Co.
for the AEC, prepared the raw plutonium fuel.
Mound Laboratory, Miamisburg, Ohio, operated by Monsanto
Research Corp., for the AEC, fabricated the raw fuel into the final fuel
form and encapsulated the fuel.
PHOTOGRAPHIC EQUIPMENT
Still and motion pictures will be made of most spacecraft maneuvers and
crew lunar surface activities, and mapping photos from orbital altitude to
aid in planning future landing missions. During lunar surface activities,
emphasis will be on photographic documentation of lunar surface features
and lunar material sample collection.
Camera equipment stowed in the Apollo 13 command module
consists of two 70mm Hasselblad electric cameras, a 16mm motion
picture camera, and the Hycon lunar topographic camera (LTC).
The LTC, to be flown on Apollos 13, 14 and 15, is stowed beneath the
commander's couch. In use, the camera mounts in the crew access
hatch window.
The LTC with 18-inch focal length f/4.0 lens provides resolution of
objects as small as 15-25 feet from a 60-nm altitude and as small as 3 to
5 feet from the 8-nm pericynthion. Film format is 4.5-inch square frames
on 100 foot long rolls, with a frame rate variable from 4 to 75 frames a
minute. Shutter speeds are 1/50, 1/100, and 1200 second. Spacecraft
forward motion during exposures is compensated for by a servo- '
controlled rocking mount. The film is held flat in the focal plane by a
vacuum platen connected to the auxiliary dump valve.
The camera weighs 65 pounds without film, is 28 inches long, 10.5
inches wide, and 12.25 inches high. It is a modification of an aerial
reconnaissance camera.
Future lunar landing sites and targets of scientific interest will be
photographed with the lunar topographic camera in overlapping sequence of
single frame modes. A candidate landing site northwest of the crater
Censorinus will be photographed from the 8-mile pericynthion during the
period between descent orbit insertion and CSM/LM separation. Additional
topographic photos of the Censorinus site and sites near Davy Rille and
Descartes will be made later in the mission from the 60-nm circular orbit.
The camera again will be unstowed and mounted for 20 minutes of
photography of the lunar disc at 5 minute intervals starting at 2 hours
after transearth injection.
Cameras stowed in the lunar module are two 70mm - Hasselblad data cameras
fitted with 60mm Zeiss Metric lenses, a 16mm motion picture camera fitted
with a 10mm lens and a Kodak closeup stereo camera for high resolution
photos on the lunar surface. The LM Hasselblads have crew chest mounts
that leave both hands free.
One of the command module Hasselblad electric cameras is normally
fitted with an 80mm f/2.8 Zeiss Planar lens, but bayonet mount
250mm and 500mm lenses may be substituted for special tasks.
The second Hasselblad camera is fitted with an 80mm lens and a
Reseau plate which allows greater dimensional control on
photographs of the lunar surface. The 500mm lens will be used only
as a backup to the lunar topographic camera.
The 80mm lens has a focusing range from 3 feet to infinity' and has a
field of view of 38 degrees vertical and horizontal on the square- format
film frame. Accessories for the command module Hasselblads include a
spotmeter, interalometer remote control cable, and film magazines.
Hasselblad shutter speeds range from time exposure and one second to one
1/500 second.
The Maurer 16mm motion picture camera in the command module has lenses of
5, 18, and 75mm available. The camera weighs 2.8 pounds with a 130-foot
film magazine attached. Accessories include a right-angle mirror, a power
cable, and a sextant adapter which allows the camera to use the navigation
sextant optical system. The LM motion picture camera will be mounted in the
right-hand window to record descent and landing and the two EVA periods
and later will be taken to the surface.
The 35 mm stereo closeup camera stowed in the LM MESA shoots 24mm square
stereo pairs with an image scale of one half actual size.  The camera is
fixed focus and is equipped with a stand-off hood to position the camera
at the proper focus distance. A long handle permits an EVA crewman to
position the camera without stooping for surface object photography.
Detail as small as 40 microns can be recorded. The camera allows
photography of significant surface structure which would remain intact
only in the lunar environment, such as fine powdery deposits, cracks or
holes, and adhesion of particles. A battery-powered electronic flash
provides illumination, and film capacity is a minimum of 100 stereo pairs.
LUNAR DESCRIPTION
Terrain - Mountainous and crater-pitted, the mountains rising as high as
29 thousand feet and the craters ranging from a few inches to 180 miles in
diameter. The craters are thought to be formed primarily by the impact of
meteorites.  The surface is covered with a layer of fine grained material
resembling silt or sand, as well as small rocks and boulders.
Environment - No air, no wind, and no moisture. The temperature ranges
from 243 degrees F. in the two-week lunar day to 279 degrees below zero in
the two-week lunar night. Gravity is one-sixth that of Earth.
Micrometeoroids pelt the Moon since there is no atmosphere to burn them
up.  Radiation might present a problem during periods of unusual solar
activity.
Far Side - The far or hidden side of the Moon no longer is a complete
mystery. It was first photographed by a Russian craft and since then
has been photographed many times, particularly from NASA's Lunar
Orbiter and Apollo - spacecraft.
Origin - There is still no agreement among scientists on the origin of
the Moon. The three theories: (1) the Moon once was part of Earth
and split off into its own orbit, (2) it evolved as a separate body at
the same time as Earth, and (3) it formed elsewhere in space and
wandered until it was captured by Earth's gravitational field.
Physical Facts
Diameter  		2,160 miles (about 1/4 that of Earth)
Circumference 		6,790 miles (about 1/4 that of Earth
Distance from Earth 	238,857 miles (mean, 221,463 minimum
 			to 252,710 maximum)
Surface temperature 	+243 degrees F (Sun at zenith) -279 degrees F (night)
Surface gravity 	1/6 that of Earth
Mass 			l/lOOth that of Earth
Volume 			1/50th that of Earth
Lunar day and night 	14 Earth days each
Mean velocity in orbit 	2,287 miles-per-hour
Escape velocity  	1.48 miles-per-second
Month (period of 	27 days, 7 hours, 43 minutes
rotation around Earth)
Landing Site
The landing site selected for Apollo 13 is located at 3 degrees 40' 7"S,
17 degrees 27' 3"W, about 30 miles north of the Fra Mauro crater. The site
is in a hilly, upland region. This will be the first Apollo landing to
other than a lunar mare, the flat dark areas of the Moon once thought to
be lunar seas. This hilly region has been designated as the Fra Mauro
formation, a widespread geological unit covering large portions of the
lunar surface around Mare Imbrium (Sea of Rains). The Fra Mauro formation
is interpreted by lunar geologists to be an ejecta blanket of material
thrown out by the event which created the circular Mare Imbrium basin.
The interpretation of the Fra Mauro formation as ejecta from Mare
Imbrium gives rise to the expectation that surface material
originated from deep within the Moon, perhaps from a hundred
miles below the Moon's surface. If the interpretation proves correct,
it will also be possible to date the Mare Imbrium event, believed to
be a major impact, perhaps the in-fall of a smaller Moon, which was
swept up in the primordial, accretionary evolution of the Moon.
Based on this theory, rocks from the Fra Mauro formation should
predate the rocks returned from either Apollo 11 (4.6 billion years)
or Apollo 12 (3.5 billion years) and be close to the origin age of the
Moon.
APOLLO 13 FLAGS, LUNAR MODULE PLAQUE
The United States flag to be erected on the lunar surface measures
30 by 48 inches and will be deployed on a two-piece aluminum tube
eight feet long. The folding horizontal bar which keeps the flag
standing out from the staff on the airless Moon has been improved
over the mechanisms used on Apollo 11 and 12.
The flag, made of nylon, will be stowed in the lunar module descent
stage modularized equipment stowage assembly (MESA) instead of in
a thermal-protective tube on the LM front leg, as in Apollo 11 and
12.
Also caroled on the mission and returned to Earth will be 25 United
States and 50 individual state flags, each 4 by 6 inches.
A 7 by 9 inch stainless steel plaque, similar to those flown on
Apollos 11 and 12, will be fixed to the LM front leg. The plaque has
on it the words "Apollo 13" with "Aquarius" beneath, the date, and
the signatures of the three crewmen.
SATURN V LAUNCH VEHICLE
The Saturn V launch vehicle (SA-508) assigned to the Apollo 13
mission was developed at the Marshall Space Flight Center,
Huntsville, Ala. The vehicle is almost identical to those used in the
missions of Apollo 8 through 12.
First Stage
The first stage (S-IC) of the Saturn V is built by the Boeing Company
at NASA's Michoud Assembly Facility, New Orleans, La. The stages
five F-l engines develop a total of about 7.6 million pounds of thrust
at launch. Major components of the stage are the forward skirt,
oxidizer tank, intertank structure, fuel tank, and thrust structure.
Propellant to the five engines normally flows at a rate of 29,364.5
pounds (3,400 gallons) each second. One engine is rigidly mounted on
the stage's centerline; the other four engines are mounted on a ring
at 90 degree angles around the center engine. These four outer engines are
gimbaled to control the vehicle's attitude during flight.
Second Stage
The second stage (S-II) is built by the Space-Division of the North
American Rockwell Corporation at Seal Beach, Calif. Five J-2 engines
develop a total of about 1.16 million pounds of thrust during light.
Major structural components are the forward skirt, liquid hydrogen
and liquid oxygen tanks (separated by an insulated common
bulkhead), a thrust structure, and an interstage section that connects
the first and second stages. The five engines are mounted and used
in the same way as the first stage's F-l engines: four outer engines
can be gimbaled; the center one is rigid.
Third Stage
The third stage (S-IVB) is built by the McDonnell Douglas
Astronautics Company at Huntington Beach, Calif. Major components
are the aft interstage and skirt, thrust structure, two propellant
tanks with a common bulkhead, a forward skirt, and a single J-2
engine. The gimbaled engine has a maximum thrust of 230,000
pounds, and can be shut off and restarted.
Instrument Unit
The instrument unit (IU), built by the International Business
Machines Corp., at Huntsville, Ala., contains navigation, guidance and
control equipment to steer the launch vehicle into its Earth orbit and
into translunar trajectory. The six major systems are structural,
thermal control, guidance and control, measuring and telemetry,
radio frequency, and electric.
The instrument unit provides a path-adaptive guidance scheme
where in a programmed trajectory is used during first stage boost
with guidance beginning during second stage burn. This scheme
prevents movements that could cause the vehicle to break up while
attempting to compensate for winds or jet streams in the
atmosphere.
The instrument unit's inertial platform (heart of the navigation,
guidance and control system) provides space-fixed reference
coordinates and measures acceleration along three mutually
perpendicular axes of a coordinate system. If the platform fails
during boost, systems in the Apollo spacecraft are programmed to
provide guidance for the launch vehicle. After second stage ignition,
the spacecraft commander could manually steer the vehicle in the
event of loss of the launch vehicle inertial platform.
Propulsion
The Saturn V has 37 propulsive units, with thrust ratings ranging
from 70 pounds to more than 1.5 million pounds.  The large main
engines burn liquid propellants; the smaller units use solid or
hypergolic propellants.
The five F-l engines on the first stage burn a combination of RP-l
(kerosene) as fuel and liquid oxygen as oxidizer. Each engine
develops approximately 1,516,918 pounds of thrust at liftoff,
building to about 1,799,022 pounds before cutoff. The five-engine
cluster gives the first stage a thrust range of from 7,584,593 pounds
at liftoff to 8,995,108 pounds just before center engine cutoff. The F-l
engine weighs almost 10 tons, is more than 18 feet long and has a
nozzle exit diameter of nearly 14 feet. The engine consumes almost
three tons of propellant every second.
The first stage also has eight solid-fuel retrorockets that fire to
separate the first and second stages.  Each retrorocket produces a
thrust of 87,900 pounds for 0.6 seconds.
The second and third stages are powered by J-2 engines that burn
liquid hydrogen (fuel) and liquid oxygen (oxidizer).
J-2 engine thrust varies from 184,841 to 232,263 pounds
during flight. The 3,500-pound J-2 engine is considered
more efficient than the F-l engine because the J-2 burns high-
energy liquid hydrogen. F-l and J-2 engines are built by
the Rocketdyne Division of the North American Rockwell Corp.
The second stage also has four 21,000-pound-thrust solid fuel ullage
rockets that settle liquid propellant in the bottom of the main tanks
and help attain a "clean' separation from the first stage. Four retrorockets, located in the S-IVB's aft
interstage (which never separates from the S-II), separate the S-II
from the S-IVB. There are two jettisonable ullage rockets for
propellant settling before engine ignition. Eight smaller engines in
the two auxiliary propulsion system modules on the S-IVB stage
provide three-axis attitude control.
COMMAND AND SERVICE MODULE STRUCTURE, SYSTEMS
The Apollo spacecraft for the Apollo 13 mission is comprised of
Command Module 109, Service Module 109, Lunar Module 7, a
spacecraft-lunar module adapter (SLA) and a launch escape system.
The SLA houses the lunar module and serves as a mating structure
between the Saturn V instrument unit and the SM.
Launch Escape System (LES) -- Would propel command module to
safety in an aborted launch. It has three solid-propellant rocket
motors: a 147,000 pound-thrust launch escape system motor, a
2,400-pound-thrust pitch control motor, and a 31,500 pound-thrust
tower jettison motor. Two canard vanes deploy to turn the command
module aerodynamically to an attitude with the heat-shield forward.
The system is 33 feet tall and 4 feet in diameter at the base, and
weighs 8,945 pounds.
Command Module (CM) Structure -- The command module is a
pressure vessel encased in heat shields, cone-shaped, weighing
12,365 pounds at launch.
The command module consists of a forward compartment which
contains two reaction control engines and components of the Earth
landing system; the crew compartment or inner pressure vessel
containing crew accommodations, controls and displays, and many of
the spacecraft systems; and the aft compartment housing ten
reaction control engines, propellant tankage, helium tanks, water
tanks, and the CSM umbilical cable. The crew compartment contains
210 cubic feet of habitable volume.
Heat-shields around the three compartments are made of brazed
stainless steel honeycomb with an outer layer of phenolic epoxy
resin as an ablative material.
CSM and LM are equipped with the probe-and-drogue docking
hardware. The probe assembly is a powered folding coupling and
impact attenuating device mounted in the CM tunnel that mates
with a conical drogue mounted in the LM docking tunnel. After the
12 automatic docking latches are checked following a docking
maneuver, both the probe and drogue are removed to allow crew
transfer between the SM and LM.
Service Module (SM) Structure -- At launch, the service module
for the Apollo 13 mission will weigh 51,105 pounds. Aluminum
honeycomb panels one inch thick form the outer skin, and milled
aluminum radial beams separate the interior into six sections around
a central cylinder containing two helium spheres, four sections
containing service propulsion system fuel-oxidizer tankage, another
containing fuel cells, cryogenic oxygen and hydrogen, and one sector
essentially empty.
Spacecraft-LM Adapter (SLA) Structure -- The spacecraft LM adapter
is a truncated cone 28 feet long tapering from 260 inches diameter
at the base to 154 inches at the forward end at the service module
mating line. The SLA weighs 4,000 pounds and houses the LM during
launch and Earth orbital flight.
CSM Systems
Guidance, Navigation and Control System (GNCS) -- Measures and
controls spacecraft position, attitude, and velocity, calculates
trajectory, controls spacecraft propulsion system thrust vector, and
displays abort data. The guidance system consists of three
subsystems: Inertial, made up of an inertial measurement unit and
associated power and data components; computer which processes
information to or from other components; and optics consisting of
scanning telescope and sextant for celestial and/or landmark sighting
for spacecraft navigation. VHF ranging device serves as a backup to
the LM rendezvous radar.
Stabilization and Control Systems (SCS) -- Controls spacecraft
rotation, translation, and thrust vector and provides displays for
crew-initiated maneuvers; backs up the guidance system for control
functions. It has three subsystems; attitude reference, attitude
control, and thrust vector control.
Service Propulsion System (SPS) -- Provides thrust for large spacecraft
velocity changes through a gimbal-mounted 20,500-pound-thrust hypergolic
engine using a nitrogen tetroxide oxidizer and a 50-50 mixture of
unsymmetrical dimethyl hydrazine and hydrazine fuel. This system is in the
service module. The system responds to automatic firing commands from the
guidance and navigation system or to manual commands from the crew. The
engine thrust level is not throttleable. The stabilization and control
system gimbals the engine to direct the thrust vector through the
spacecraft center of gravity.
Telecommunications System -- Provides voice, television, telemetry, and
command data and tracking and ranging between the spacecraft and Earth,
between the command module and the lunar module and between the spacecraft
and astronauts during EVA. It also provides intercommunications between
astronauts.
The high-gain steerable S-Band antenna consists of four, 31 inch diameter
parabolic dishes mounted on a folding boom at the aft end of the service
module. Signals from the ground stations can be tracked either
automatically or manually with the antenna's gimballing system. Normal
S-Band voice and uplink/downlink communications will be handled by the
omni and high-gain antennas.
Sequential System -- Interfaces with other spacecraft systems and
subsystems to initiate time critical functions during launch, docking
maneuvers, sub-orbital aborts, and entry portions of a mission. The
system also controls routine spacecraft sequencing such as service
module separation and deployment of the Earth landing system.
Emergency Detection System (EDS) -- Detects and displays to the
crew launch vehicle emergency conditions, such as excessive pitch or
roll rates or two engines out, and automatically or manually shuts
down the booster and activates the launch escape system; functions
until the spacecraft is in orbit.
Earth Landing System (ELS) -- Includes the drogue and main
parachute system as well as post-landing recovery aids. In a normal
entry descent, the command module forward heat shield is
jettisoned at 24,000 feet, permitting mortar deployment of two
reefed 16.5-foot diameter drogue parachutes for orienting and
decelerating the spacecraft. After disreef and drogue release, three
mortar deployed pilot chutes pull out the three main 833 foot
diameter parachutes with two-stage reefing to provide gradual
inflation in three steps. Two main parachutes out of three can
provide a safe landing.
Reaction Control System (RCS) -- The SM RCS has four identical RCS
"quads" mounted around the SM 90 degrees apart. Each quad has
four 100 pound-thrust engines, two fuel and two oxidizer tanks and
a helium pressurization sphere. Attitude control and small velocity
maneuvers are made with the SM RCS.
The CM RCS consists of two independent six-engine subsystems of six
93 pound-thrust engines each used for spacecraft attitude control
during entry. Propellants for both CM and SM RCS are monomethyl
hydrazine fuel and nitrogen tetroxide oxidizer with helium
pressurization. These propellants burn spontaneously when
combined (without an igniter).
Electrical Power System (EPS) -- Provides electrical energy sources, power
generation and control, power conversion, conditioning, and distribution
to the spacecraft. The primary source of electrical power is the fuel
cells mounted in the SM. The fuel cell also furnishes drinking water to
the astronauts as a by-product.
Three silver-zinc oxide storage batteries supply power to the CM
during entry and after landing, provide power for sequence
controllers, and supplement the fuel cells during periods of peak
power demand. A battery charger assures a full charge prior to
entry.
Two other silver-zinc oxide batteries supply power for explosive devices
for CM/SM separation, parachute deployment and separation, third-stage
separation, launch escape tower separation, and other pyrotechnic uses.
Environmental Control System (ECS) -- Controls spacecraft atmosphere,
pressure, and temperature and manages water. In addition to regulating
cabin and suit gas pressure, temperature and humidity, the system removes
carbon dioxide, odors and particles and ventilates the cabin after
landing. It collects and stores fuel cell potable water for crew use,
supplies water to the glycol evaporators for cooling, and dumps surplus
water overboard through the waste H O dump nozzle. Proper operating
temperature of electronics an electrical equipment is maintained by this
system through the use of the cabin heat exchangers, the space radiators,
and the glycol evaporators.
Recovery Aids -- Recovery aids include the uprighting system,
swimmer interphone connections, sea dye marker, flashing beacon,
VHF recovery beacon, and VHF transceiver. The uprighting system
consists of three compressor-inflated bags to upright the spacecraft if
it should land in the water apex down (stable II position).
Caution and Warning System -- Monitors spacecraft systems for
out-of-tolerance conditions and alerts crew by visual and audible alarms.
Controls and Displays -- Provide status readouts and control
functions of spacecraft systems in the command and service modules.
All controls are designed to be operated by crewmen in pressurized
suits. Displays are grouped by system and located according to the
frequency of use and crew responsibility.
LUNAR MODULE STRUCTURES WEIGHT
The lunar module is a two-stage vehicle designed for space
operations near and on the Moon. The lunar module stands 22 feet
11 inches high and is 31 feet wide (diagonally across landing gear).
The ascent and descent stages of the LM operate as a unit until
staging, when the ascent stage functions as a single spacecraft for
rendezvous and docking with the CM.
Ascent Stage
Three main sections make up the ascent stage: the crew compartment,
midsection, and aft equipment bay. Only the crew compartment and
midsection are pressurized (4.8 psig). The cabin volume is 235 cubic feet
(6.7 cubic meters). The stage measures 12 feet 4 inches high by 14 feet 1
inch in diameter. The ascent stage has six substructural areas: crew
compartment, midsection, aft equipment bay, thrust chamber assembly
cluster supports, antenna supports, and thermal and micrometeoroid shield.
The cylindrical crew compartment is 92 inches (2.35 m) in diameter and 42
inches (l.07 m) deep. Two flight stations are equipped with control and
display panels, armrests, body restraints, landing aids, two front windows,
an overhead docking window, and an alignment optical telescope in the
center between the two flight stations. The habitable volume is 160 cubic
feet.
A tunnel ring atop the ascent stage meshes with the command module docking
latch assemblies. During docking, the CM docking ring and latches are
aligned by the LM drogue and the CSM probe.
The docking tunnel extends downward into the midsection 16 inches (40 cm).
The tunnel is 32 inches (81 cm) in diameter and is used for crew transfer
between the CSM and LM. The upper hatch on the inboard end of the docking
tunnel opens inward and cannot be opened without equalizing pressure on
both hatch surfaces.
A thermal and micrometeoroid shield of multiple layers of Mylar and
a single thickness of thin aluminum skin encases the entire ascent
stage structure.
Descent Stage
The descent stage center compartment houses the descent engine,
and descent propellant tanks are housed in the four square bays
around the engine.  Quadrant II (Seq bay) contains ALSEP, and
Radioisotope Thermoelectric Generator (RTG) externally. Quadrant IV
contains the MESA. The descent stage measures 10 feet 7 inches high
by 14 feet 1 inch in diameter and is encased in the Mylar and
aluminum alloy thermal and micrometeoroid shield.
The LM egress platform, or "porch", is mounted on the forward
outrigger just below the forward hatch. A ladder extends down the
forward landing gear strut from the porch for crew lunar surface
operations.
The landing gear struts are explosively extended and provide lunar
surface landing impact attenuation. The main struts are filled with
crushable aluminum honeycomb for absorbing compression loads.
Footpads 37 inches (0.95 m) in diameter at the end of each landing
gear provide vehicle support on the lunar surface.
Each pad (except forward pad) is fitted with a 68 inch long lunar
surface sensing probe which signals the crew to shut down the
descent engine upon contact with the lunar surface.
LM-7 flown on the Apollo 13 mission has a launch weight of 3,476
pounds. The weight breakdown is as follows:
Ascent stage, dry    	  4,668 lbs.   	Includes water
			          	and oxygen, no crew
Descent stage, dry 	  4,650 lbs.
RCS propellants (loaded)    590 lbs.
DPS propellants (loaded) 18,339 lbs.
APS propellants (loaded)  5,229 lbs.
			 33,476 lbs.
Lunar Module Systems
Electrical Power System -- The LM DC electrical system consists of
six silver zinc primary batteries -- four in the descent stage and two
in the ascent stage. Twenty-eight-volt DC power is distributed to all
LM systems. AC power (117v 400 Hz) is supplied by two inverters.
Environment Control System - Consists of the atmosphere revitalization
section of oxygen supply and cabin pressure control section, water
management, heat transport section, and outlets for oxygen and water
reservicing of the portable life support system (PLSS).
Components of the atmosphere revitalization section are the suit
circuit assembly which cools and ventilates the pressure garments,
reduces carbon dioxide levels, removes odors, noxious gases
and excessive moisture; the cabin recirculation assembly which
ventilates and controls cabin atmosphere temperatures; and the
steam flex duct which vents to space steam from the suit circuit
water evaporator.
The oxygen supply and cabin pressure section supplies gaseous
oxygen to the atmosphere revitalization section for maintaining suit
and cabin pressure. The descent stage oxygen supply provides
descent flight phase and lunar stay oxygen needs, and the ascent
stage oxygen supply provides oxygen needs for the ascent and
rendezvous flight phase.
Water for drinking cooling, fire fighting, food preparation, and refilling
the PLSS cooling water servicing tank is supplied by - the water
management section. The water is contained in three nitrogen-pressurized
bladder-type tanks, one of 367-pound capacity in the descent stage and two
of 47.5-pound capacity in the ascent stage.
The heat transport section has primary and secondary water glycol
solution coolant loops. The primary coolant loop circulates water-
glycol for temperature control of cabin and suit circuit oxygen and
or thermal control of batteries and electronic components mounted
on cold plates and rails. If the primary loop becomes inoperative, the
secondary loop circulates coolant through the rails and cold plates
only. Suit circuit cooling during secondary coolant loop operation is
provided by the suit loop water boiler. Waste heat from both loops is
vented overboard by water evaporation or sublimators.
Communications System -- Two S-band transmitter-receivers, two
VHF transmitter-receivers, a signal processing assembly, and
associated spacecraft antenna make up the LM communications
system. The system transmits and receives voice and tracking and
ranging data, and transmits telemetry data on about 270
measurements and TV signals to the ground. Voice communications
between the LM and ground stations is by S-band, and between the
LM and CSM voice is on VHF.
Although no real-time commands can be sent to the LM, the digital
uplink processes guidance officer commands, such as state vector
updates, transmitted from Mission Control Center to the LM
guidance computer.
The data storage electronics assembly (DSEA) is a four-channel voice
recorder with timing signals, with a 10-hour recording capacity,
which will be brought back into the CSM for return to Earth. DSEA
recordings cannot be "dumped" to ground stations.
LM antennas are one 26-inch-diameter parabolic S-band steerable
antenna, two S-band inflight antennas, two VHF inflight antennas,
EVA antenna, and an erectable S-band antenna (optional) for lunar
surface.
Guidance, Navigation, and Control System -- Comprised of six
sections: primary guidance and navigation section (PGNS), abort
guidance section (AGS), radar section, control electronics section
(CES), and orbit rate display Earth and lunar (ORDEAL).
The PGNS is an aided inertial guidance system updated by the alignment
optical telescope, an inertial measurement unit, and the rendezvous and
landing radars. The system provides inertial reference data for
computations, produces inertial alignment reference by feeding optical
sighting data into the LM guidance computer, displays position and
velocity data, computes LM-CSM rendezvous data from radar inputs, controls
attitude and thrust to maintain desired LM trajectory, and controls
descent engine throttling and gimbaling.
The LM-7 primary guidance computer has the Luminary IC Software
program, which is an improved version over that in LM-6.
The AGS is an independent backup system for the PGNS, having its
own inertial sensors and computer.
The radar section is made up of the rendezvous radar which provides CSM
range and range rate, and line-of-sight angles for maneuver computation to
the LM guidance computer; and the landing radar which provides altitude
and velocity to the M guidance computer during lunar landing. The
rendezvous radar has an operating range from 80 feet to 400 nautical
miles. The ranging tone transfer assembly, utilizing VHF electronics, is a
passive responder to the CSM VHF ranging device and is a backup to the
rendezvous radar.
The CES controls LM attitude and translation about all axes. It also
controls by PGNS command the automatic operation of the ascent and
descent engine and the reaction control thruster. Manual attitude
controller and thrust-translation controller commands are also
handled by the CES.
ORDEAL, displayed on the flight director attitude indicator, is the
computed local vertical in the pitch axis during circular Earth or lunar
orbits.
Reaction Control System -- The LM has four RCS engine clusters of four
100-pound (45.4 kg) thrust engines each, which use helium pressurized
hypergolic propellants. The oxidizer is nitrogen tetroxide, fuel is
Aerozine (50/50 blend of hydrazine and unsymmetrical dimethyl hydrazine).
Interconnect valves permit the RCS system to draw from ascent engine
propellant tanks.
The RCS provides small stabilizing impulses during ascent and
descent burns, controls LM attitude during maneuvers, and produces
thrust for separation, and for ascent/descent engine tank ullage.
The system may be operated in either the pulse or steady-state
modes.
Descent Propulsion System -- Maximum rated thrust of the descent
engine is 9,870 pounds (4,380.9 kg) and is throttleable between
1,050 pounds (476.7 kg) and 6,300 pounds (2,860 2 kg). The engine
can be gimbaled six degrees in any direction in response to attitude
commands and to compensate for center of gravity offsets.
Propellants are helium-pressurized Aerozine 50 and nitrogen
tetroxide.
Ascent Propulsion System -- The 3,500-pound (1,589 kg) thrust
ascent engine is not gimbaled and performs at full thrust. The engine
remains dormant until after the ascent stage is separated from the
descent stage. Propellants are the same as are burned by the RCS
engines and the descent engine.
Caution and Warning, Controls and Displays -- These two systems
have the same function aboard the lunar module as they do aboard
the command module (See CSM systems section.)
Tracking and Docking Lights -- A flashing tracking light (once per
second, 20 milliseconds duration) on the front face of the lunar
module is an aid for contingency CSM-active rendezvous LM rescue.
Visibility ranges from 400 nautical miles through the CSM sextant to
130 miles with the naked eye. Five docking lights analogous to
aircraft running lights are mounted on the LM for CSM-active
rendezvous: two forward yellow lights, aft white light, port red light
and starboard green light. All docking lights have about a l,000-foot
visibility.
APOLLO 13 CREW AND CREW EQUIPMENT
Life Support Equipment - Space Suits
Apollo 13 crewmen will wear two versions of the Apollo space suit: an
intravehicular pressure garment assembly worn by the command module pilot
and the extravehicular pressure garment assembly worn by the commander and
the lunar module pilot.  Both versions are basically identical except that
the extravehicular version has an integral thermal/meteoroid garment over
the basic suit.
From the skin out, the basic pressure garment consists
of a comfort layer, a neoprene-coated nylon pressure' bladder and a nylon
restraint layer.  The outer layers of the intravehicular suit are,
from the inside out, nomex and two layers of elon-coated Beta cloth. The
extravehicular integral thermal/meteoroid cover consists of a liner of two
layers of neoprene-coated nylon, seven layers of Beta/Kapton spacer
laminate, and an outer layer of Teflon-coated Beta fabric.
The extravehicular suit, together with a liquid cooling garment, portable
life support system (PLSS), oxygen purge system, lunar extravehicular
visor assembly and other components make up the extravehicular mobility
unit (EMU). The EMU provides an extravehicular crewman with lie support for
an hour mission outside the lunar module without replenishing
expendables.  EMU total weight is 183 pounds. The intra-vehicular suit
weighs 35.6 pounds.
Liquid cooling garment--A knitted nylon-spandex garment with
tubing through which cooling water - from the PLSS is circulated. It is
worn next to the skin and replaces the constant wear-garment during EVA
only.
Portable life support system -- A backpack supplying oxygen at 3.9 psi and
cooling water to the liquid cooling garment.  Return oxygen is cleansed of
solid and gas contaminants by a lithium hydroxide canister. The PLSS
includes communications and telemetry equipment, displays and controls,
and a main power supply. The PLSS is covered by a thermal insulation
jacket. (Two stowed in LM).
Oxygen purge system--Mounted atop the PLSS, the oxygen purge
system provides a contingency 45-minute supply of gaseous oxygen
in two two-pound bottles pressurized to 5,880 psia. The system may
also be worn separately on the front of the pressure garment
assembly torso. It serves as a mount for the YHF antenna for the
PLSS. (Two stowed in LM).
Lunar extravehicular visor assembly--A polycarbonate shell and two visors
with thermal control and optical coatings on them. The EVA visor is
attached over the pressure helmet to provide impact, micrometeoroid,
thermal and ultraviolet-infrared light protection to the EVA crewmen.
Since Apollo 12, a sunshade has been added to the outer portion of the
LEVVA in the middle potion of the helmet rim.
Extravehicular gloves--Built of an outer shell of Chromel-R fabric and
thermal insulation to provide protection when handling extremely
hot and cold objects. The finger tips are made of silicone rubber to
provide more sensitivity.
A one-piece constant-wear garment, similar to "long johns, is worn
as an undergarment for the space suit in intravehicular operations
and for the inflight coveralls. The garment is porous-knit cotton with
a waist-to-neck zipper for donning. Biomedical harness attach points
are provided.
During periods out of the space suits, crewmen wear two-piece Teflon
fabric inflight coveralls for warmth and for pocket stowage of
personal items.
Communications carriers ("Snoopy Hats") with redundant
microphones and earphones are worn with the pressure helmet; a
lightweight headset is worn with the inflight coveralls.
Another modification since Apollo 12 has been the addition of eight-ounce
drinking water bags ("Gunga Dins" attached to the inside neck rings of the
EVA suits. The crewmen can take a sip of water from the 6 X 8 inch bag
through a 1/8-inch diameter tube within reach of his mouth. The bags are
filled from the lunar module potable water dispenser.
Apollo Lunar Hand Tools
Special Environmental Container - The special environmental sample is
collected in a carefully selected area and sealed in a special container
which will retain a high vacuum. The container is opened in the Lunar
Receiving Laboratory where it will provide scientists the opportunity to
study lunar material in its original environment.
 Extension handle - This tool is of aluminum alloy tubing with a malleable
stainless steel cap designed to be used as an anvil surface.  The handle
is designed to be used as an extension for several other tools and to
permit their use without requiring the astronaut to kneel or bend down.
The handle is approximately 24 inches long and 1 inch in diameter. The
handle contains the female half of a quick disconnect fitting designed to
resist compression, tension, torsion, or a combination of these loads.
Three core tubes - These tubes are designed to be driven or augered
into loose gravel, sandy material, or into soft rock such as feather
rock or pumice. They are about 15 inches in length and one inch in
diameter and are made of aluminum tubing. Each tube is supplied
with a removable non-serrated cutting edge and a screw-on cap
incorporating a metal-to-metal crush seal which replaces the cutting
edge. The upper end of each tube is sealed and designed to be used
with the extension handle or as an anvil. Incorporated into each tube
is a spring device to retain loose materials in the tube.
Scoops (large and small) - These tools are designed for use as a
trowel and as a chisel. The scoop is fabricated primarily of aluminum
with a hardened-steel cutting edge riveted on and a nine-inch
handle. A malleable stainless steel anvil is on-the end of the handle.
The angle between the scoop pan and the handle allows a
compromise for the dual use. The scoop is used either by itself or
with the extension handle.  The large scoop has a sieve which permits
particles smaller than 1/2 cm to pass through.
Sampling hammer - This tool serves three functions, as a sampling
hammer, as a pick or mattock, and as a hammer to drive the core
tubes or scoop. The head has a small hammer face on one end, a
broad horizontal blade on the other, and large hammering flats on
the sides. The handle is 14 inches long and is made of formed tubular
aluminum. The hammer has on its lower end a quick-disconnect to
allow attachment to the extension handle for use as a hoe. The head
weight has been increased to provide more impact force.
The tongs are designed to allow the astronaut to retrieve small samples
from the lunar surface while in a standing position.  The tines are of
such angles, length, and number to allow samples of from 3/8 up to 2-1/2-
inch diameter to be picked up. This tool is 24 inches in overall length.
Brush/Scriber/Hand Lens - A composite tool
(1) Brush - To clean samples prior to selection
(2) Scriber - To scratch samples for selection and to mark for
identification
(3) Hand lens - Magnifying glass to facilitate sample selection
Spring Scale - To weigh two rock boxes and other bags containing
lunar material samples, to maintain weight budget for return to
Earth.
Instrument staff - The staff hold the Hasselblad camera. The staff
breaks down into sections. The upper section telescopes to allow
generation of a vertical stereoscopic base of one foot for
photography. Positive stops are provided at the extreme of travel. A
shaped hand grip aids in aiming and carrying. The bottom section is
available in several lengths to suit the staff to astronauts of varying
sizes. The device is fabricated from tubular aluminum.
Gnomon - This tool consists of a weighted staff suspended on a two-
ring gimbal and supported by a tripod. The staff extends 12 inches
above the gimbal and is painted with a gray scale. The gnomon is
used as a photographic reference to indicate local vertical, sun
angle, and scale. The gnomon has a required accuracy of vertical
indication of 20 minutes of arc. Magnetic damping is incorporated to
reduce oscillations.
Color Chart - The color chart is painted with three primary colors and a
gray scale. It is used as a calibration for lunar photography. The scale
is mounted on the tool carrier but may easily be removed and returned to
Earth for reference. The color chart is 6 inches in size.
Carrier - The carrier is the stowage container for the lunar flight. After
the landing the carrier serves as support for the astronaut when he kneels
down, as a support for the sample bags and samples, and as a tripod base
for the instrument staff.  The carrier folds flat for stowage. For field
use it opens into a triangular configuration. The carrier is constructed
of formed sheet metal and approximates a truss structure. Six-inch legs
extend from the carrier to elevate the carrying handle sufficiently to be
easily grasped by the astronaut.
Field Sample Bags - Approximately 80 bags four inches by five
inches are included in the Apollo lunar hand tools for the packaging
of samples. These bags are fabricated from Teflon FEP.
 Collection Bag - This is a large bag (4 X 8 inches) attached to the
astronaut's side of the tool carrier. Field sample bags are stowed in this
bag after they have been filled. It can also be used for general storage
or to hold items temporarily.  (Two in each SRC).
Trenching Tool - A trenching tool with a pivoting scoop has been provided
for digging the two-foot deep soil mechanics investigation trench. The
two-piece handle is five feet long. The scoop is eight inches long and
five inches wide and pivots from in-line with the handle to 90
degrees-similar to the trenching tool carried on infantry backpacks. The
trenching tool is stowed in the MESA rather than in the tool carrier.
Lunar Surface Drill - The 29.4-pound Apollo Lunar Surface Drill
(ALSD) is stowed in the ALSEP subpackage No. 2 and will be used for
boring two ten-foot deep 1.25-inch diameter holes for ALSEP heat
flow experiment probes, and one approximately eight-foot-deep,
one-inch-diameter core sample. The silver zinc battery-powered
rotary percussive drill-has a clutch to limit torque to 20 foot-pounds.
A treadle assembly serves as a drilling platform and as a core stem
lock during the drill string decoupling operation as the string is
withdrawn from the lunar soil. Bore stems for the heat low
experiment holes are of boron/fiberglass and the core sample core-
stems are titanium. Cutting bits are tungsten carbide.
Apollo 13 Crew Menu-
More than 70 items comprise the food selection list of freeze-
dried-rehydratable, wet-pack and spoon-bowl foods. Balanced
meals for five days have been packed in man/day overlaps. Items
similar to those in the daily menus have been packed in a snack
pantry. The snack pantry permits the crew to locate easily a food
item in a smorgasbord mode without having to find a regular meal
somewhere down deep in a storage box
Water for drinking and rehydrating food is obtained from two sources in
the command module--a dispenser for drinking water and a water-spigot at
the food preparation station supplying water at about 155 or 55 degrees F.
The portable water dispenser squirts water continuously as long as the
trigger is held down and the food preparation spigot dispenses water in
one-ounce increments.
A continuous-feed hand water dispenser similar to the one in the
command module is used aboard the lunar module for cold-water
rehydration of food packets stowed aboard the LM.
After water has been injected into a food bag, it is kneaded or
about three minutes. The bag neck is then cut off and the food
squeezed into the crewman's mouth. After a meal, germicide pills
attached to the outside of the food bas are placed in the bags to
prevent fermentation and gas formation. The bags are then rolled
and stowed in waste disposal compartments.
The day-by-day, meal-by-meal Apollo 13 Menu for Commander Lovell is on the
following page as a typical everyday menu for each crewman.
Personal Hygiene
Crew personal hygiene equipment aboard Apollo 13 includes body
cleanliness items, the waste management system and one medical
kit.
Packaged with the food are a toothbrush and a two-ounce tube of
toothpaste for each crewman. Each man-meal package
contains a 3.5-by-4-inch wet-wipe cleansing towel. Additionally,
three packages of 12-by-12-inch dry towels are stowed
beneath the command module pilot's couch. Each package contains
seven towels. Also stowed under the command module pilot's couch
are seven tissue dispensers containing 53 three-ply tissues each.
Solid body wastes are collected in plastic defecation bags which
contain a germicide to prevent bacteria and gas formation. The bags
are sealed after use and stowed in empty food containers for post-
flight analysis.
Urine collection devices are provided for use while wearing either
the pressure suit or the inflight coveralls. The urine is dumped
overboard through the spacecraft urine dump valve in the CM and
stored in the LM.
Medical Kit
The 5x5x8-inch medical accessory kit is stowed in a compartment on the
spacecraft right side wall beside the lunar module pilot couch. The medical
kit contains three motion sickness injectors three pain suppression
injectors, one two ounce bottle first aid ointment, two one-ounce bottles
eye drops, three nasal sprays, two compress bandages, 12 adhesive
bandages, one oral thermometer, and four spare crew biomedical harnesses.
Pills in the medical kit are 60 antibiotic, 12 nausea, 12 stimulant, 18
pain killer, 60 decongestant, 24 diarrhea, 72 aspirin and 21 sleeping.
Additionally, a small medical kit containing four stimulant, eight
diarrhea, two sleeping and four pain killer pills, 12 aspirin, one bottle
eye drops, two compress bandages, 8 decongestant pills, one pain injector,
one bottle nasal spray is stowed in the lunar module flight data file
compartment.
Survival Gear
The survival kit is stowed in two rucksacks in the right-hand forward
equipment bay above the lunar module pilot.
Contents of rucksack No. l are: two combination survival lights, one
desalter kit, three pair sunglasses, one radio beacon, one spare radio
beacon battery and spacecraft connector cable, one knife in sheath,
three water containers, and two containers of Sun lotion, two utility-
knives, three survival blankets and one utility netting.
Rucksack No. 2: one three-man life raft with C02 inflater, one sea
anchor two sea dye markers, three sunbonnets, one mooring
lanyard, three mainlines and two attach brackets.
The survival kit is designed to provide a 48-hour postlanding (water or
land) survival capability for three crewmen between 40 degrees North and
South latitudes.
Biomedical Inflight Monitoring
The Apollo 13 crew biomedical telemetry data received by the Manned Space
Flight Network will be relayed for instantaneous display at Mission
Control Center where heart rate and breathing rate data will be displayed
on the flight surgeon's console. Heart rate and respiration rate average
range and deviation are computed and displayed on digital TV screens.
In addition, the instantaneous heart rate, real-time and delayed EKG
and respiration are recorded on strip charts for each man.
Biomedical telemetry will be simultaneous from all crewmen while
in the CSM, but selectable by a manual onboard switch in the LM.
Biomedical data observed by the flight surgeon and his team in the
Life Support Systems Staff Support Room will be correlated with
spacecraft and space suit environmental data displays.
Blood pressures are no longer telemetered as they were in the
Mercury and Gemini programs. Oral temperatures, however, can be
measured onboard for diagnostic purposes and voiced down by the
crew in case of inflight illness.
Energy expended by the crewmen during-EVA will be determined indirectly
using a metabolic computation program based on three separate measurements
1) Heart rate portion -- Heart rate will be determined from telemetered EK
and converted to oxygen consumption (litre/min) and heat production BTU
hour based on preflight calibration curves. These curves are determined
from exercise response tests utilizing a bicycle ergometer.
2) Oxygen usage portion -- Oxygen usage will be determined from the
telemetered measurement of PSS oxygen supply pressure.  Suit leak
determined pre-flight is taken into account. Heat production will be
calculated from oxygen usage
3) Liquid cooled garment temperature portion -- The amount of heat-taken
up by the liquid cooled garment will be determined from telemetered
measurements of the LCG water temperature inlet and change in/out.  This
measurement (the amount of heat taken up-by the water), plus an allowance
made for sensible and latent heat loss radiant heat load, and possible
heat storage will provide an indication Or heat production by the
crewman.
Training
The crewmen of Apollo 13 have spent more than five hours of formal
crew training for each hour of the lunar-launching mission's ten-day
duration. More than 1,000 hours o training were in Apollo 13 crew
training syllabus over and above the normal preparations for the
mission--technical briefings and reviews, pilot meetings and study.
The Apollo 13 crewmen also took part in prelaunch testing at
Kennedy Space Center, such as altitude chamber tests and the
countdown demonstration tests (CDDT) which provided the crew with
thorough operational knowledge of the complex vehicle.
Highlights of specialized Apollo 13 crew training topics are:
* Detailed series of briefings on spacecraft systems, operation and
modifications.
* Saturn launch vehicle briefings on countdown, range safety, flight
dynamics, failure modes and abort conditions. The launch vehicle
briefings were updated periodically.
* Apollo Guidance and Navigation system briefings at the
Massachusetts Institute of Technology Instrumentation Laboratory.
* Briefings and continuous training on mission photographic
objectives and use of camera equipment.
* Extensive pilot participation in reviews of all flight procedures for
normal as well as emergency situations.
* Stowage reviews and practice in training sessions in the spacecraft,
mockups and command module simulators allowed the crewmen to evaluate
spacecraft stowage of crew-associated equipment.
* More than 400 hours of training per man in command module and
lunar module simulators at MSC and KSC, including closed-loop
simulations with flight controllers in the Mission Control Center.
Other Apollo simulators at various locations were used extensively
for specialized crew training.
* Lunar surface briefings and some 20 suited l-g walk throughs of lunar
surface EVA operations covering lunar geology and microbiology and
deployment of experiments in the Apollo Lunar Surface Experiment Package
(ALSEP).  Training in lunar surface EVA included practice sessions with
lunar surface sample gathering tools and return containers, cameras, the
erectable S-band antenna and the modular equipment stowage assembly:(MESA)
housed-in the LM descent stage.' '
* Proficiency flights in the lunar landing training vehicle (LLTV) for
the commander.
* Zero-g and one-sixth g aircraft flights using command module an
lunar module mockups for EVA and pressure suit doffing/donning
practice and training.
* Underwater zero-g training in the MSC Water Immersion Facility using
spacecraft mockups to further familiarize crew with all aspects of GSM-LM
docking tunnel intravehicular transfer and EVA in pressurized suits.
* Water egress training conducted in indoor tanks as well as in the Gulf
of Mexico, included uprighting from the Stable II position (apex down)
to the Stable I position (apex up), egress onto rafts donning Biological
Isolation Garments ('BIGs), decontamination procedures and helicopter
pickup.
* Launch pad egress training from mockups and from the actual
spacecraft on the launch pad for possible emergencies such as fire,
contaminants and power failures.
* The training covered use of Apollo spacecraft fire
suppression equipment in the cockpit.
* Planetarium reviews at Morehead Planetarium Chapel Hill, N.C., and at
Griffith Planetarium, Los Angeles, Calif., of the celestial sphere with
special emphasis on the 37 navigational stars used by the Apollo guidance
computer.
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
WASHINGTON, D. C.
BIOGRAPHICAL DATA
NAME: James Arthur Lovell, Jr. Captain, USN
NASA Astronaut.
BIRTHPLACE AND DATE: Born March 25, 19:28, in Cleveland, Ohio.
 His mother, Mrs. Blanche Lovell, resides at Edgewater Beach, Florida.
PHYSICAL DESCRIPTION:: Blond hair, blue eyes; height: 5 feet: 11
inches, weight: 170 pounds.
EDUCATION: Graduated from Juneau High School, Milwaukee, Wisconsin;
attended the University of Wisconsin for 2 years, then received a Bachelor
of Science degree from the United States Naval Academy in 1952, presented
an Honorary Doctorate from Illinois Wesleyan University in 1969.
MARITAL STATUS: Married to the-former Marilyn Gerlach of Milwaukee,
Wisconsin. Her parents, Mr. and Mrs. Carl Gerlach, are residents of
Milwaukee.
CHILDREN: Barbara L., October 13, 1953; James A., February :15,
1955, Susan K., July 14, 1958; Jeffrey C.,
January 14, 1966.
RECREATIONAL INTERESTS: His hobbies are golf, swimming,
handball, and tennis.
ORGANIZATIONS: Member of the Society of Experimental Test Pilots
and the Explorers Club.
SPECIAL HONORS: Awarded-the NASA Distinguished Service Medal, two NASA
Exceptional Service Medals, the Navy Astronaut Wings, the Navy
Distinguished Service Medal, and two Navy Distinguished Flying Crosses,
recipient of the 1967 FAI De Laval and Gold Space Medals (Athens, Greece),
the American Academy of Achievement Golden late Award, the City of New
York Gold Medal in 1969, the City of Houston Medal for Valor in 1969, the
National Geographic Society's Hubbard Medal in 1969, the National Academy
Of Television Arts and Sciences Special Trustees Award in 1969, and the
Institute of Navigation Award in 1969.
Co-recipient of the American Astronautical Society Flight
Achievement Awards in 1966 and 1968, the Harmon International
Trophy in 1966 and 1967, the Robert H. Goddard Memorial Trophy in
1969 the H. H. Arnold Trophy or 1969, the General Thomas D.
White USAF Space Trophy for 1968, the Robert J. Collier Trophy for
1968, and the 1969 Henry Bennett Distinguished Service Award.
EXPERIENCE: Lovell, a Navy Captain, received flight training following
graduation from Annapolis in 1952.
He has had numerous naval aviator assignments including a 4-year tour as a
test pilot at the Naval Air Test Center, Patuxent River, Maryland. While
there he served as program manager for the F4H weapon system evaluation. A
graduate of the Aviation Safety School of the University of Southern
California, he also served as a light instructor and safety engineer with
Fighter Squadron 101 at the Naval Air Station, Oceana, Virginia.
He has logged more than 4,407 hours flying time--more than 3,000
hours in jet aircraft.
CURRENT -ASSIGNMENT: Captain Lovell was selected as an astronaut
by NASA in September 1962. He has since served as backup pilot for
the Gemini 4 flight and backup command pilot for the Gemini 9
flight.
On December 4, 1965, he and Command pilot Frank Borman were
launched into space on the history-making Gemini 7 mission. The
flight lasted 330 hours and 35 minutes, during which the following
space firsts were accomplished: longest manned space flight; first
rendezvous of two manned maneuverable spacecraft, as Gemini 7
was Joined in orbit by Gemini 6; and longest multi-manned space
flight. It was also on this flight that numerous technical and medical
experiments were completed successfully.
The Gemini 12 mission, with Lovell and pilot Edwin Aldrin, began
on November 11, 1966. This 4-day, 59-revolution flight brought the
Gemini Program to a successful close. Major accomplishments of
the 94-hour 35-minute flight included a third-revolution
rendezvous with the previously launched Agena (using for the first
time backup onboard computations due to radar failure); a tethered
station keeping exercise; retrieval of a micrometeorite experiment
package from the spacecraft exterior; an evaluation of the use of
body restraints specially designed for completing work tasks outside
of the spacecraft; and completion of numerous photographic
experiments, highlights of which are the first pictures taken from
space of an eclipse of the sun.
Gemini 12 ended with retrofire at the beginning of the 60th
revolution, followed by the second consecutive fully automatic
controlled reentry of a spacecraft, and a landing in the Atlantic
within 2 1/2 miles of the USS WASP.
As a result of his participation in the Gemini 7 and 12 flights, Lovell
logged 425 hours and 10 minutes in space. Aldrin established a new
EVA record by completing 5 1/2 hours outside the spacecraft during
two standup EVAs and one umbilical EVA.
Lovell served as command module pilot for the epic six day journey
of Apollo 8--man's maiden voyage to the moon --December 21-27,
1968. Apollo 8 was the first manned spacecraft to be lifted into
near-earth orbit by a 7 1/2-million pound thrust Saturn V launch
vehicle, and all events in the flight plan occurred as scheduled with
unbelievable accuracy.
A "go"' for the translunar injection burn was given midway through the
second near-earth orbit, and the restart of the S-IVB third stage to
effect this maneuver increased the spacecraft's velocity to place it on an
intercept course with the moon. Lovell and fellow crew members, Frank
Borman (spacecraft commander) and William A. Anders (lunar module pilot),
piloted their spacecraft some 223,000 miles to become the first humans to
leave the earth's influence; and upon reaching the moon on December 24,
they performed the first critical maneuver to place Apollo 8 into a 60 by
168 nautical miles lunar orbit.
Two revolutions later, the crew executed a second maneuver using the
spacecraft's 20,500-pound thrust service module propulsion system to
achieve a circular lunar orbit of 60 nautical miles. During their ten
revolutions of the moon, the crew conducted live television transmissions
of the lunar surface and performed such tasks as landmark and Apollo
landing site tracking, vertical stereo photography and stereo navigation
photography, and sextant navigation using lunar landmarks and stars. At
the end of the tenth lunar orbit, they executed a transearth injection
burn which placed Apollo 8 on a proper trajectory for the return to earth.
The final leg of the trip required only 58 hours as compared to the 69
hours used to travel to the moon, and Apollo 8 came to a successful
conclusion on December 27, 1968. Splashdown occurred at an estimated 5,000
yards from the USS YORKTOWN, following the successful negotiation of a
critical 28-mile high reentry corridor at speeds close to 25,000 miles per
hour.
Captain Lovell has since served as the backup spacecraft commander
for the Apollo 11 lunar landing mission. He has completed three
space flights and holds the U.S. Astronaut record for time in space
with a total of 572 hours and 10 minutes.
SPECIAL ASSIGNMENT: In addition to his regular duties as an
astronaut, Captain Lovell continues to serve as Special Consultant to
the President's Council on Physical Fitness and Sports--an assignment
he has held since June 1967.
CURRENT SALARY: $1,717.28 per month.
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
BIOGRAPHICAL DATA
NAME: Thomas-Kenneth Mattingly II Lieutenant Commander, USN)
NASA-Astronaut
BIRTHPLACE AND DATE: Born in Chicago, Ill., March 17, 1936. His
parents, Mr. and Mrs. Thomas K. Mattingly, now reside in Hialeah,
Fla.
PHYSICAL DESCRIPTION: Brown hair; blue eyes; height: 5 feet  10
inches; weight: 140 pounds.
EDUCATION: Attended Florida elementary and secondary schools and is a
graduate of Edison High School, Miami, Fla.; received a Bachelor of
Science degree in Aeronautical Engineering from Auburn University in 1958.
MARITAL STATUS: Single
RECREATIONAL INTERESTS: Enjoys water skiing and playing handball and
tennis.
ORGANIZATIONS: Member of the American Institute of Aeronautics; and
Astronautics and the U.S. Naval Institute.
EXPERIENCE: Prior to reporting for duty at the Manned Spacecraft Center,
he was a student at the Air Force Aerospace Research Pilot School.
He began his Naval career as an Ensign in 1958 and received his wings in
1960. He was then assigned to VA-35 and flew AlH aircraft aboard the USS
SARATOGA from 1960 to 1963. In July 1963, he served in VAH deployed
aboard the USS Franklin D. ROOSEVELT where he flew the A3B aircraft for
two years.
He has logged 31O0 hours of flight time--1,946 hours
in jet aircraft.
CURRENT ASSIGNMENT:  Commander Mattingly is one of the 19 astronauts
selected by NASA in April 1966. He served as a member of the astronaut
support crews for the Apollo 8 and 11 missions.
CURRENT SALARY: 1,293.33 per month.
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
WASHINGTON, D. C.
Biographical DATA
NAME: Fred Wallace Haise, Jr. (Mr.)
NASA Astronaut
BIRTHPLACE AND DATE: Born in Biloxi, Miss., on Nov. 14, 1933; his
mother, Mrs. Fred W. Haise, Sr., resides in Biloxi.
PHYSICAL DESCRIPTION: Brown hair; brown eyes; height: 5 feet 9 1/2
inches; weight: 150 pounds.
EDUCATION: Graduated from Biloxi High School, Biloxi, Miss.;
attended Perkinston Junior College (Association of Arts); received a
Bachelor of Science degree with honors in Aeronautical Engineering
from the University of Oklahoma in 1959.
MARITAL STATUS: Married to the former Mary Griffin Grant of
Biloxi, Miss. Her parents, Mr. and Mrs. William J. Grant Jr., reside in
Biloxi.
CHILDREN: Mary M., January 25, 1956; Frederick T., May 13, 1958;
Stephen W., June 30, 1961.
ORGANIZATIONS: Member of the Society of Experimental Test Pilots,
Tau Beta Pi, Sigma Gamma Tau, and Phi Theta Kappa.
SPECIAL HONORS: Recipient of the A.B. Honts Trophy as the
outstanding graduate of class 64A from the Aerospace Research Pilot
School in 1964, awarded the American Defense Ribbon and the
Society of Experimental Test Pilots Ray E. Tenhoff Award for 1966.
EXPERIENCE: Haise was a research pilot at the NASA Flight Research Center
at Edwards, Calif., before coming to Houston and the Manned Spacecraft
Center; and from September 1959 to March 1963, he was a research pilot at
the NASA Lewis Research Center in Cleveland, Ohio. During this time, he
authored the following papers which have been published: a NASA TND,
entitled "An Evaluation of the Flying Qualities of Seven General-Aviation
Aircraft;" NASA TND 3380, "Use of Aircraft for Zero Gravity Environment,
May 1966;" SAE Business Aircraft Conference Paper, entitled "An Evaluation
of General Aviation Aircraft Flying Qualities," March 30-April 1, 1966;
and a paper delivered at the tenth symposium of the Society of
Experimental Test Pilots, entitled "A Quantitative/Qualitative Handling
Qualities Evaluation of Seven General-Aviation Aircraft,"  1966.
He was the Aerospace Research Pilots School's outstanding graduate of
Class 64A and served with the U.S. Air Force from October 1961 to August
1962 as a tactical fighter pilot and as Chair of the 164th
Standardization-Evaluation Flight of the 164th Tactical Fighter Squadron
at Mansfield Ohio. from March 1957 to September 1959, he was a
fighter-interceptor pilot with the 185th Fighter Interceptor Squadron in
the Oklahoma Air National Guard.
He also served as a tactics and all weather flight instructor in the U.S.
Navy Advanced Training Command at NAAS Kingsville, Texas, and
was assigned as a U.S. Marine Corps fighter pilot to VMF-533 and
114 at MCAS Cherry Point, N.C., from March 1954 to September
1956.
His military career began in October 1952 as a Naval Aviation Cadet
at the Naval Air Station in Pensacola, Fla.
He has accumulated 5,800 hours flying time, including 3,000 hours in
jets.
CURRENT ASSIGNMENT: Mr. Haise is one of the 19 astronauts selected
by NASA in April 1966. He served as backup lunar module pilot for
the Apollo 8 and 11 missions.
CURRENT SALARY: $1,698.00 per month.
LAUNCH COMPLEX 39
Launch Complex 39 facilities at the Kennedy Space Center were
planned and built specifically for the Apollo Saturn V, the space
vehicle being used in the United States manned lunar exploration
program.
Complex 39 introduced the mobile concept of launch operations in
which the space vehicle is thoroughly checked out in an enclosed
building before it is moved to the launch pad for final preparations.
This affords greater protection from the elements and permits a high
launch rate since pad time is minimal.
Saturn V stages are shipped to the Kennedy Space Center by
ocean-going vessels and specially designed aircraft. Apollo spacecraft
modules are transported by air and first taken to the Manned
Spacecraft Operations Building in the Industrial Area south of
Complex 39 for preliminary checkout, altitude-chamber testing, and
assembly.
Apollo 12 is the sixth-Saturn V/Apollo space vehicle to be launched from
Complex 39's Pad A, one of two octagonal launch pads which are 3,000 feet
across. The major components of Complex 39 include:
1. The Vehicle Assembly Building, heart of the complex, is where the
363-foot-tall space vehicle is assembled and tested. It contains 129.5
million cubic feet of space, covers eight acres, is 716 feet long and 518
feet wide. Its high bay area, 525 feet high, contains four assembly and
checkout bays and its low bay area - 210 feet high 42 feet wide and 274
feet long - contains eight stage-preparation and checkout cells. There are
141 lifting devices in the building, ranging from one-ton hoists to two
250-ton high lift bridge cranes.
2. The Launch Control Center, a four-story structure adjacent and to
the south of the Vehicle Assembly Building is a radical departure
from the dome-shaped, "hardened" blockhouse at older launch sites.
The Launch Control Center is the electronic "brain" of Complex 39 and
was used for checkout and test operations while Apollo 12 was being
assembled inside the Vehicle Assembly Building high bay. Three of
the four firing rooms contain identical sets of control and monitoring
equipment so that launch of one vehicle and checkout of others may
continue simultaneously. Each firing room is associated with a ground
computer facility to provide data links with the launch vehicle on its
mobile launcher at the pad or inside the Vehicle Assembly
building.
3. The Mobile Launcher, 445 feet tall and weighing 12 million
pounds, is a transportable launch base and umbilical tower for the
space vehicle.
4. The Transporters, used to move mobile launchers into the Vehicle
Assembly Building and then - with their space vehicles - to the
launch pad, weigh six million pounds and are among the largest
tracked vehicles known. The Transporters -there are two - are 131
feet long and 114 feet wide. Powered by electric motors driven by
two 2,750-horsepower diesel engines, the vehicles move on four
double-tracked crawlers, each 10 feet high and 40 feet long.
Maximum speed is about one-mile-per-hour loaded and two miles-
per-hour unloaded. The three and one-half mile trip to Pad A with a
mobile launcher and space vehicle takes approximately seven hours.
Apollo 12 rollout to the pad occurred on December 15, 1969.
5. The Crawlerway is the roadway for the transporter and is 131 feet
wide divided by a median strip. This is the approximate width of an
eight-lane turnpike and the roadbed is designed to accommodate a
combined weight of more than 18 million pounds.
6. The Mobile Service Structure is a 402-foot-tall, 9.8 million pound
tower used to service the Apollo space vehicle at the pad. Moved into
place about the Saturn V/Apollo space vehicle and its mobile
launcher by a transporter, it contains five work platforms and
provides 360-degree platform access to the vehicle being prepared
for launch. It is removed to a parking area about 11 hours before
launch.
7. A Water Deluge System will provide about a million gallons of
industrial water for cooling and fire prevention during the launch of
Apollo 13. The water is used to cool the mobile launcher, the flame
trench and the flame deflector above which the mobile launcher is
positioned.
8. The Flame Deflector is an "A"-shaped, 1.3 million pound structure
moved into the flame trench beneath the launcher prior to launch. It
is covered with a refractory material designed to withstand the
launch environment. The flame trench itself is 58 feet wide and
approximately six feet above mean sea level at the base.
9. The Pad - A and B - are octagonal in shape and have center hardstands
constructed of heavily reinforced concrete. The top of Pad A stands about
48 feet above sea level. Saturn V propellants - liquid oxygen liquid
hydrogen and RP-l, the latter a high grade kerosene - are stored in large
tanks spaced near the pad perimeter and carried by pipelines from the
tanks to the pad, up the mobile launcher and into the launch vehicle
propellant tanks. Also located in the pad area are pneumatic, high
pressure gas, electrical, and industrial water support activities. Pad B,
used for the launch of Apollo 10, is located 8,700 feet north of Pad A.
MISSION CONTROL CENTER
The Mission Control Center at the Manned Spacecraft Center,
Houston, is the focal point for Apollo flight control activities. The
center receives tracking and telemetry data from the Manned Space
Flight Network which in turn is processed by the MCC Real-Time
Computer Complex or display to flight controllers in the Mission
Operations Control Room (MOCR) and adjacent staff support rooms.
Console positions in the two identical MOCRs in Mission Control
Center fall into three basic operations groups: mission command and
control, systems operations and flight dynamics.
Positions in the command and control group are:
Mission Director -- responsible for overall mission conduct.
Plight Operations Director -- represents MSC management.
Flight Director -- responsible for operational decisions and actions
in the MOCR.
Assistant Flight Director -- assists flight director and acts in-hi
absence.
Flight Activities Officer -- develops and coordinates flight plan.
Department of Defense Representative -- coordinates and directs
DOD mission support.
Network Controller -- responsible to FD for Manned Space Flight
Network status and troubleshooting; MCC equipment operation.
Surgeon -- monitors crew medical condition and informs FD of any
medical situation affecting mission.
 Spacecraft Communicator (Capcom) -- serves as voice contact with
flight .
 Experiments Officer -- coordinates operation and control of onboard
flight experiments.
 Public Affairs Officer -- reports mission progress to public through
commentary and relay of live air-to-ground transmissions.
Systems Operations Group:
Environmental, Electrical and Instrumentation Engineer (EECOM) --
monitors and troubleshoots command/service module
environmental, electrical, and sequential systems.
Guidance, Navigation and Control Engineer (GNC) --monitors and
troubleshoots CSM guidance, navigation, control, and propulsion
systems.
 LM Environmental and Electrical Engineer (TELCOM) --LM
counterpart to EECOM.
 LM Guidance, Navigation and Control Engineer (Control)--LM
counterpart to GNC.
Booster Systems Engineer (BSE) (three positions) --responsible for
monitoring launch vehicle Performance and for sending function
commands.
Communications Systems Engineer (CSE) (call sign INCO) and
Operations and Procedures Officer (O&P) -- share responsibility for
monitoring and troubleshooting spacecraft and lunar surface
communication systems and for coordinating MCC procedures with
other NASA centers and the network.
Flight Dynamics Group:
* Flight Dynamics Officer (FIDO) -- monitors towered flight events
and plans spacecraft maneuvers.
* Retrofire Officer (Retro) -- responsible for deorbit maneuvers in Earth
orbit and entry calculation on lunar return trajectories.
* Guidance Officer (Guido) -- responsible for monitoring and updating CSM
and LM guidance systems and for monitoring systems performance during
powered flight.
Each MOCR operations group has a staff support room on the same floor in
which detailed monitoring and analysis is  conducted. Other supporting MCC
areas include the spaceflight Meteorological Room, the Space Environment
(radiation) Console, Spacecraft Planning and Analysis (SPAN) Room for
detailed spacecraft performance analysis, Recovery-Operations Control Room
and the Apollo Lunar Surface Experiment Package Support Room.
Located on the first floor of the MCC are the communications, command, and
telemetry system (CCATS) for processing incoming data from the tracking
network, and the real-time computer complex (RTCC) which converts
flight data into displays usable to MOCR light controllers.
MANNED SPACE FLIGHT NETWORK
 The worldwide Manned Space light Network MSFN provides reliable,
continuous, and instantaneous communications with the astronauts, launch
vehicle, and spacecraft from liftoff to splashdown.  Following the flight,
the network will continue in support of the link between Earth and the
Apollo experiments left on the lunar surface by the Apollo crew.
 The MSFN is maintained and operated by the NASA Goddard Space Flight
Center Greenbelt, Md., under the direction of NASA's Office of Tracking
and Data Acquisition. In the MSFN Operations Center (MSFNOC) at Goddard,
the Network Director and his term of Operations Managers, with the
assistance of a Network support Team, keep the entire complex tuned for the
mission support. Should Houston's mission control center be seriously
impaired for an extended time, the Goddard Center becomes an emergency
mission control center.
The MSN employs 12 ground tracking stations equipped with 30- and
85-foot-antennas, an instrumented tracking ship, and four
instrumented-aircraft. For-Apollo 13, the network will be augmented by
the 210-foot antenna systems at Goldstone, Calif. and at Parkes,
Australia, (Australian Commonwealth' Scientific and Industrial Research
Organization).
NASA Communications Network (NASCOM). The tracking network
is linked together by the NASA Communications Network. All
information flows to and from MCC Houston and the Apollo
spacecraft over this communications system.
The NASCOM consists of almost three million circuit miles of densely
routed communications channels. It uses satellites, submarine cables,
land lines, microwave systems, and high frequency radio facilities for
access links.
NASCOM control center is located at Goddard.   Regional
communication switching centers are in London, Madrid, Canbera,
Australia, Honolulu and Guam.
Three Intelsat communications satellites will be used for Apollo 13.
One satellite over the Atlantic will link Goddard with stations at
Madrid, Canary Islands, Ascension and the Vanguard tracking ship.
Another Atlantic satellite will provide a direct link between Madrid
and Goddard for TV signals received from the spacecraft. The third
satellite over the mid-Pacific will link Carnarvon, Canberra, and
Hawaii with Goddard through a ground station at Brewster Flats,
Wash.
At Goddard, NASCOM switching computers simultaneously send the
voice signals directly to the Houston flight controllers and
the tracking and telemetry data to computer processing complexes
at Houston and Goddard. The Goddard Real Time Computing Complex
verifies performance of the tracking network and uses the collected
tracking data to drive displays in the Goddard Operations Control
Center.
Establishing the Link -- The Merritt Island tracking station monitors
prelaunch test, the terminal countdown, and the first minutes of
launch.
An Apollo instrumentation ship (USNS VANGUARD) fills the gaps
beyond the range of land tracking stations. For Apollo 13 this ship
will be stationed in the Atlantic to cover the insertion into Earth
orbit. Apollo instrumented aircraft provide communications support
to the land tracking stations during translunar injection and reentry
and cover a selected abort area in the event of "no-go" decision after
insertion into Earth orbit.
Lunar Bound - Approximately one hour after the spacecraft has been
injected into its translunar trajectory (some 10,000 miles from the
Earth), three prime tracking stations spaced nearly equidistant around the
Earth will take over tracking and communicating with Apollo.
Each of the prime stations, located at Goldstone, Madrid and
Canberra, has a dual system for use when tracking the command
module in lunar orbit and the lunar module in separate flight paths
or at rest on the Moon. These stations are equipped with 85-foot
antennas.
The Return Tri -- To make an accurate reentry, data from the tracking
stations are fed into the MCC computers to develop necessary information
for the Apollo 13 crew.
Appropriate MSFN stations, including the aircraft in the
Pacific, provide support during the reentry.
Through the journey to the Moon and return, television will be received
from the spacecraft at the three prime stations. In addition, a 210-foot
antenna at Goldstone (an antenna of NASA's Deep Space Network) will
augment the television coverage while Apollo 13 is near and on the Moon.
For black and white TV, scan converters at the stations permit immediate
transmission of commercial quality TV via NASCOM to Houston, where it will
be released to U.S. TV networks.
Black and white TV can be released simultaneously in Europe and
the Far East through the MSFN stations in Spain and Australia.
For color TV, the signal will be converted to commercial
quality at the MSC Houston. A black and white version of the color
signal can be released locally simultaneously through the stations in
Spain and Australia.
Network Computers
At fraction-of-a-second intervals, the network's digital data
processing systems, with NASA's Manned Spacecraft Center as the
focal point, "talk" to each other or to the spacecraft. High-speed
computers at the remote sites (tracking ship included) relay
commands or "up-link" data on such matters as control of cabin
pressure, orbital guidance commands, or "go-no-go" indications to
perform certain functions.
When information originates from Houston, the computers refer to
their pre-programmed information for validity before transmitting
the required data to the spacecraft.
Such "up-link" information is communicated at a rate of about 1,200
bits-per-second. Communication of spacecraft data between remote
ground sites and the Mission Control Center, via high-speed
communications links, occurs at twice the rate. Houston reads
information from these ground sites at 8,800 bits-per-second.
The computer systems perform many other functions, including:
Assuring the quality of the transmission lines by continually testing
data paths.
Verifying accuracy of the messages.
Constantly updating the flight status.
For "down-link" data, sensors built into the spacecraft continually sample
cabin temperature, pressure and physical information on the astronauts such
as heartbeat and respiration. These data are transmitted to the ground
stations at 51.2 kilobits (12,800 decimal digits) per second.
At MCC the computers:
Detect and select changes or deviations, compare with their stored
programs, and indicate the problem areas or pertinent data to the
flight controllers;
Provide displays to mission personnel;
Assemble output data in proper formats;
Log data on magnetic tape for the flight controllers.
The Apollo Ship Vanguard
The USNS Vanguard will perform tracking, telemetry, and
communication functions for the launch phase and Earth orbit
insertion. Vanguard will be stationed about 1,000 miles southeast of
Bermuda (28 degrees N., 49 degrees W.).
Apollo Range Instrumentation Aircraft (ARI)
During the Apollo 13 TLI maneuver, two ARIA will record telemetry
data from Apollo and relay voice communication between the
astronauts and the Mission Control Center at Houston. The ARIA will
be located between Australia and Hawaii.
For reentry, two ARIA will be deployed to the landing area to relay
communications between Apollo and Mission Control at Houston and
provide position information on the spacecraft after the blackout
phase of reentry has passed.
The total ARIA fleet for APOLLO missions consists of four EC-135A Boeing
707) jets with 7-foot parabolic antennas installed in the nose section.
CONTAMINATION CONTROL PROGRAM
In 1966 an Interagency Committee on Back Contamination (ICBC) was
established to assist NASA in developing a program to prevent
contamination of the Earth from lunar materials following manned
lunar exploration and to review and approve plans and procedures to
prevent back contamination. Committee membership includes
representatives from Public Health Service, Department of
Agriculture, Department of the Interior, NASA, and the National
Academy of Sciences.
The Apollo Back Contamination Program can be divided into three
phases. The first phase covers procedures which are followed by the
crew while in flight to reduce and, if possible, eliminate the return of
lunar surface contaminations in the command module.
The second phase includes recovery, isolation, and transport of the
crew, spacecraft, and lunar samples to the Manned Spacecraft Center.
The third phase encompasses quarantine operations and preliminary
sample analysis in the Lunar Receiving Laboratory.
A primary step in preventing back contamination is careful attention
to spacecraft cleanliness following lunar surface operations. This
includes use of special cleaning equipment, stowage provisions for
lunar-exposed equipment, and crew procedures for proper
"housekeeping.'
Prior to reentering the LM after lunar surface exploration, the
crewmen brush lunar surface dust or dirt from the space suit using
special brushes. They will scrape their overboots on the LM footpad
and while ascending the LM ladder, dislodge any clinging particles by
a kicking action.
After entering and pressurizing the LM cabin, the crew doff their
portable life support system, oxygen purge system, lunar boots, EVA
gloves, etc.
Following LM rendezvous and docking with the CM, the CM tunnel
will be pressurized and checks made to ensure that an adequate
pressurized seal has been made. During the period, some of the
equipment may be vacuumed.
The lunar module cabin atmosphere will be circulated through the
environmental control system suit circuit lithium hydroxide (LiOH)
canister to filter particles from the atmosphere. A minimum of five
hours weightless operation and filtering will essentially eliminate the
original airborne particles.
The CM pilot will transfer lunar surface equipment stowage bags into
the LM one at a time. The equipment transferred will be bagged
before being transferred.  The only equipment which will not be
bagged at this time are the crewmen's space suits and flight logs.
Command Module Operations - Through the use of operational and
housekeeping procedures the command module cabin will be purged
of lunar surface and/or other particulate contamination prior to
Earth reentry. These procedures start while the LM is docked with
the CM and continue through reentry into the Earth's atmosphere.
During subsequent lunar orbital flight and the transearth phase, the
command module atmosphere will be continually filtered through
the environmental control system lithium hydroxide canister. This
will remove essentially all airborne dust particles. After about 96
hours operation essentially none of the original contaminates will
remain.
Lunar Mission Recovery Operations
Following landing and the attachment of the flotation collar to the
command module, a swimmer will open the spacecraft hatch, pass in
three clean flight coveralls and three filter masks and close the
hatch.
Crew retrieval will be accomplished by helicopter to the carrier and
subsequent crew transfer-to the Mobile Quarantine facility. The
spacecraft will be retrieved by the aircraft carrier and isolated.
LUNAR RECEIVING LABORATORY (LRL)
The final phase of the back contamination program is completed in
the MSC Lunar receiving Laboratory. The crew and spacecraft are
quarantined for a minimum of 21 days after completion of lunar EVA
operations and are released based upon the completion of prescribed
test requirements and results. The lunar sample will be quarantined
or a period of 50 to 80 days-depending upon results of extensive
biological tests.
The LRL serves four basic purposes:
Quarantine of crew and spacecraft, the containment of lunar and
lunar-exposed materials, and quarantine testing to search for
adverse effects of lunar material upon terrestrial life.
- The preservation and protection of the lunar samples.
- The performance of time critical investigations.
- The preliminary examination of returned samples to assist in an
intelligent distribution of samples to principal investigators.
The LRL has the only vacuum system in the world with space gloves
operated by a man leading directly into a vacuum chamber at
pressures of about 10 billionth of an atmosphere. It has a low level
counting facility, the background count is an order of magnitude
better than other known counters. Additionally, it is a facility that
can handle a large variety of biological specimens inside Class III
biological cabinets designed to contain extremely hazardous
pathogenic material.
The LRL covers 83,000 square feet of floor space and includes a Crew
Reception Area (CRA), Vacuum Laboratory, Sample Laboratories
(Physical and Bio-Science) and an administrative and support area.
Special-building systems are employed to maintain air flow into
sample handling areas and the CRA, to sterilize liquid waste, and to
incinerate contaminated air from the primary containment systems.
The biomedical laboratories provide for quarantine tests to
determine the effect of lunar samples on terrestrial life. These tests
are designed to provide data upon which to base the decision to
release-lunar material from quarantine.
Among the tests:
a. Lunar material will be applied to 12 different culture media and
maintained under several environmental conditions. The media will
be observed for bacterial or fungal growth. Detailed inventories of
the microbial flora of the spacecraft and crew have been maintained
so that any living material found in the sample testing can be
compared against this list of potential contaminants taken to the
Moon by the crew or spacecraft.
b. Six types of human and animal tissue culture cell lines will be
maintained in the laboratory and together with embryonated eggs
are exposed to the lunar material. Based on cellular and/or
other changes, the presence of material can be established so
that special tests can be conducted to identify and isolate the type
of virus present.
c. Thirty-three species of plants and seedlings will be exposed to
lunar material. Seed germination, growth of plant cells or the health
of seedlings are then observed, and histological, microbiological and
biochemical techniques are used to determine the cause of any
suspected abnormality.
d.  A number of lower animals will be exposed to lunar material
including germ-free mice, fish, birds, oysters, shrimp, cockroaches,
houseflies, planaria, paramecia and euglena. If abnormalities are noted,
further tests will be conducted to determine if the condition is
transmissible from one group to another.
The crew reception area provides biological containment for the
flight crew and 12 support personnel.  The nominal occupancy is
about 14 days but the facility is designed and equipped to operate
for considerably longer.
Sterilization and Release of the Spacecraft
Postflight testing and inspection of the spacecraft is presently limited
to investigation of anomalies which happened during the flight. Generally,
this entails some specific testing of the spacecraft and removal of
certain components of systems for further analysis. The timing of
postflight testing is important so that corrective action may be taken for
subsequent flights.
The schedule calls or the spacecraft to be returned to port where a team
will deactivate pyrotechnics, and flush and drain fluid systems (except
water). This operation will be confined to the exterior of the spacecraft.
The spacecraft will then be flown to the LRL and placed in a special room
for storage, sterilization, and postflight checkout.
LUNAR RECEIVING LABORATORY TENTATIVE SCHEDULE
April 20 Activate secondary barrier; support people enter Crew Reception
Area and Central Status Station manned, LRL on mission status.
April 21 Command module landing, recovery.
April 22 First sample return container (SRC) arrives.
April 23 First SRC opened in vacuum lab, second SRC arrives; film,
tapes, LM tape recorder begin decontamination; second SRC
opened in Bioprep lab.
April 24 First sample to Radiation Counting Laboratory.
April 26 Core tube moves from vacuum lab to Physical Chemical
Lab.
April 26 MQF arrives; contingency sample goes to Physical Chemical
Lab; rock description begin in vacuum lab.
April 27 Biosample rocks move from vacuum lab to Bioprep
1ab; core tube prepared for biosample.
April 28 Spacecraft arrives.
April 29 Biosample compounded, thin-section chips sterilized out to
Thin-Section Lab, remaining samples from Bioprep Lab canned.
May 1 Thin-section preparation complete, biosample prep complete, transfer
to Physical-Chemical Lab complete, Bioprep Lab cleanup complete.
May 3 Biological protocols, Physical-Chemical Lab rock description
begin.
May 8 Crew released from CRA
May 26 Rock description complete, Preliminary Examination Team
data from Radiation Counting Lab and Gas Analysis-Lab complete.
May 28 PET data write-up and sample catalog preparation begin.
May 30 Data summary for Lunar Sample Analyses Planning Team
(LSAPT) complete.
June I LSAPT arrives.
June 2 LSAPT briefed on PET data, sample packaging begins.
June 6 Sample distribution plan complete, first batch monopole samples
canned.
June 8 Monopole experiment begins.
June 10 Initial release of Apollo 13 samples; spacecraft release.
June 14 Spacecraft equipment released
SCHEDULE FOR TRANSPORT OF SAMPLES, SPACECRAFT AND CREW
Samples
The first Apollo 13 sample return container (SRC) will be flown by
helicopter-from the deck of the USS Iwo Jima to Christmas Island, from
where it will be flown by C-130 aircraft to Hawaii. The SRC, half the
mission onboard film and any medical samples ready at the time of
helicopter departure from the Iwo Jima, will be transferred to an ARIA
(Apollo Range Instrumented Aircraft) at Hawaii for the flight to Ellington
AFB, six miles north of the Manned Spacecraft Center, with an estimated
time of arrival at 11:30 a.m. EST April 22.
The second SRC and remainder of onboard film and medical samples
will follow a similar sequence of flights the following day and will
arrive at Ellington AFB at an estimated time of 1 a.m. EST April 23.
The SRCs will be moved by auto from Ellington AFB to the Lunar
Receiving Laboratory.
Spacecraft
The spacecraft should be aboard the Iwo Jima about two hours after crew
recovery. The ship will arrive in Hawaii at 2 pm EST April 25 and the
spacecraft will be offloaded and transferred after deactivation to an
aircraft for the flight to Ellington AFB, arriving April 28. The
spacecraft will be trucked to the Lunar Receiving Laboratory where it will
enter quarantine.
Crew
The flight crew is expected to enter the Mobile Quarantine Facility (MQF)
on the Iwo Jima about 90 minutes after splashdown. Upon arrival at Hawaii,
the MQF will be offloaded and placed aboard a C-141 aircraft for the
flight to Ellington AFB, arriving at 1 am EST April 25. A transporter
truck will move the MQF from Ellington AFB to the Lunar Receiving
Laboratory--about a two-hour trip.
APOLLO PROGRAM MANAGEMENT
The Apollo Program is the responsibility of the Office of Manned
Space Flight (OMSF), National Aeronautics and Space Administration,
Washington, D.C. Dale D. Myers is Associate Administrator for
Manned Space Flight.
NASA Manned Spacecraft Center (MSC), Houston, is responsible for
development of the Apollo spacecraft, flight crew training, and flight
control. Dr. Robert R. Gilruth is Center Director.
NASA Marshall Space Flight Center (MSFC),Huntsville, Ala., is
responsible for development of the Saturn launch vehicles. Dr.
Eberhard Rees is Center Director.
NASA John F. Kennedy Space Center (KSC), Fla., is responsible for
Apollo/Saturn launch operations. Dr. Kurt H. Debus is Center Director.
The NASA Office of Tracking and Data Acquisition (OTDA) directs the
program of tracking and data flow on Apollo. Gerald M. Truszynski is
Associate Administrator for Tracking and Data Acquisition.
NASA Goddard Space Flight Center (GSFC), Greenbelt, Md., manages
the Manned Space Flight Network and Communications Network. Dr.
John F. Clark is Center Director.
The Department of Defense is supporting ASA in Apollo 13 during
launch, tracking and recovery operations. The Air Force Eastern Test
Range is responsible for range activities during launch and down-
range tracking. Recovery operations include the use of recovery ships
and Navy and Air Force aircraft.
Apollo/Saturn Officials
NASA Headquarters
Dr. Rocco A. Petrone Apollo Program Director, OMSF
Chester M. Lee (Capt., USN, Ret.) Apollo Mission Director, OMSF
Col. Thomas H. McMullen'(USAF) Apollo Assistant Mission
Director, OMSF
John D. Stevenson (Ma. Gen.,  Director of Mission Operations,
USAF, Ret.)OMSF
Ma. Gen. James W. Humphreys, Jr. Director of Space Medicine, OMSF
(USAF, MC)
John K. Holcomb,(Capt., USN, Ret.) Director of Apollo
Operations, OMSF
Lee R. Scherer,(Capt., USN, Ret.) Director of Apollo Lunar
 Exploration, OMSF
James C. Bavely Chief of Network Operations
 Branch, OTDA
Marshall Space Flight Center
Lee B. James  - Director, Program Management
Dr. F. A. Speer Manager, Mission Operations Office
Roy E. Godfrey Manager, Saturn Program Office
Matthew W. Urlaub Manager, S-IC Stage, Saturn
 Program Office
William F. LaHatte Manager, S-II Stage, Saturn
 Program Office'
Charles H. Meyers Manager (Acting), S-IVB Stage,
 Saturn Program Office
Frederich Duerr Manager, Instrument Unit,
 Saturn Program Office
William D. Brown Manager, Engine Program Office
Kennedy Space Center
Walter J. Kapryan Director of Launch Operations
Raymond L. Clark Director of Technical Support
Edward R. Mathews Apollo Program Manager
Dr. Hans F. Gruene Director, Launch Vehicle
Operations
John J. Wllliams Director, Spacecraft Operations
Paul C. Donnelly Launch Operations Manager
Isom A. Rigell Deputy Director for Engineering
Manned Spacecraft Center
Col. James A. McDivitt, (USAF) Manager, Apollo Spacecraft
  - Program
Donald K. Slayton Director, Flight Crew
 Operations
Sigurd A. Szoberg Director, Plight Operations
Milton L. Windler Flight Director
Gerald Grlffin Flight Director
Glynn S. Lunney Flight Director
Eugene F. Kranz Flight Director
Dr. Charles A. Berry Director, Medical Research and Operations
Goddard Space Flight Center
Ozro M. Covington Director of Manned Flight
 Support
William P. Varson Chief, Manned Flight Planning
 & Analysis Division
H. William Wood Chief, Manned Flight Operations
 Division
Tecwyn Roberts Chief, Manned Flight Engineering
 Division
L. R. Stelter Chief NASA Communications
Division.
Department of Defense
M. Gen. David M. Jones, (USAF) DOD Manager of Manned Space.  Flight
Support Operations, Commander of USAF Eastern -Test Range
Rear Ad Wm. S. Guest, (USN) Deputy DOD Manager of Manned- Space Flight
Support Operations,  Commander Task Force 140, - Atlantic Recovery Area
Rear Adm. Donald. C. Davis, (USN) Commander Task Force 130,
Pacific Recovery Area
Col. Kenneth J. Mask, (USAF)Director-of: W D Manned Space
  Flight Support Office
Maj. Gen. Allison C. Brooks, Commander Aerospace Rescue
and (SAF) Recovery Service
Major Apollo Saturn V Contractors
Contractor Item
Bellcomm Apollo Systems Engineering
Washington, D. C.
The Boeing Co. Technical Integration and
Washington, D. C. Evaluation
General Electric-Apollo-Systems Apollo Checkout, and Quality and
Daytona Beach, Fla Reliability
North American Rockwell Corp. Command and Service Modules
Space Div., Downey, Calif..
Grumman Aircraft Engineering Lunar Module
Corp., Bethpage, N.Y.
Massachusetts Institute of Guidance & Navigation
Technology, Cambridge, Mass.(Technical Management)
General Motors Corp., AC Guidance & Navigation
Electronics Div., Milwaukee, Wis.(Manufacturing)
TRW Inc. Trajectory Analysis
System Group LM Descent Engine
Redondo Beach, Calif. LM Abort Guidance System
Avco Corp., Space Systems Heat Shield Ablative Material
Div., Lowell, Mass.
North American Rockwell Corp. J-2 Engines, F-l Engines
Rocketdyne Div.
Canoga Park, Calif.
The Boeing Co. First Stage (SIC) of Saturn V
New Orleans, Launch Vehicles, Saturn V
Systems Engineering and Integration, Ground Support Equipment
North American Rockwell Corp. Development and Production of
Space Div. Saturn V Second Stage (S-II)
Seal Beach, Calif.
McDonnell Douglas Astronautics Development and Production of
Co., Huntington Beach, Calif. Saturn V Third Stage (S-IVB)
International Business -Machines Instrument Unit Federal Systems
Div. Huntsville, Ala.
Benix:Corp - ' Guidance-
Federal Electric Corp. Communications and Instrumentation Support,-RSC
Bendix Field Engineering Corp. Launch Operations/Complex
Support, KSC  . .
Catalytic-Dow  - Facilities Engineering and
 Modifications, KSC
- Hamilton Standard Division Portable Life Support System;
 United Aircraft Corp.  LM ECS'-
 Windsor Locks, Conn. '.
ILC Industries 'Space Suits
Dover:, Del. : '
Radio Corp. of America  - Computer - Saturn Checkout
Van.-Nubs, Calif.' '
 Sanders Associates Operational Display Systems
Nashua,  H. -  Saturn '- '
Brown Engineering.: Discrete Controls
Huntsville, Ala.
Reynolds, Smith- and Hill Engineering Design of Mobile
 Jacksonville, Fla.  Launchers
 Ingots Iron Works Mobile Launchers (ML)
 Birmingham, Ala.  '(Structural Work)
  Smith/Ernst (Joint Venture  Electrical Mechanical Portion
  Tampa, Fla. ' of MLs
  Washington, D. C.
Power Shovel, Inc.  Transporter
Marion, Ohio
  Hayes International  Mobile Launcher Service Arms
  Birmingham, Ala.
  Bendix Aerospace Systems  Apollo Surface Experiment-Package (ALSEP)
  Ann Arbor, Mich.
  Aerojet-Gen. Corp.  Service Propulsion System Engine