REPORT OF - APOLLO 13 REVIEW BOARD

REPORT OF - APOLLO 13 REVIEW BOARD
.
REPORT OF
- APOLLO 13 REVIEW BOARD
-
APPENDIX
A
BASELINE DATA: APOLLO 13
l
. FLIGHT SYSTEMS AND OPERATIONS
- VATIONAL
AERONAUTICS
AND
SPACE ADMINISTRATION
c
CONTENTS
Part
Al
A2
A2.1
Page
APPENDIX A - BASELINE DATA: APOLLO 13 FLIGHT
SYSTEMS AND OPERATIONS . . . . . . . . . . . . . .
A-l
APOLLO SPACECRAFT CONFIGURATION
A-3
..........
LAUNCH ESCAPE ASSEMBLY ..............
A-3
COMMANDMODULE..................
A-3
SERVICE MODULE ...................
A-8
SPACECRAFT LM ADAPTER ..............
A-9
SYSTEMS DESCRIPTION DATA ..............
A-11
INTRODUCTION ...................
A-11
GUIDANCE AND CONTROL ...............
A-12
Guidance
and Control
Attitude
Reference
Attitude
Control
Thrust
and Thrust
Systems
. . . . .
A-12
. . . . . . . . . . . . . . .
A-12
. . . . . . . . . . . . . . . .
A-14
. . ......
A-16
A2.2
GUIDANCE AND NAVIGATION SYSTEM (G&N) . ......
A-18
A2.3
STABILIZATION
A-20
A2.4
SERVICE PROPULSION SYSTEM (SPS)
A2.5
REACTION CONTROL SYSTEM (RCS).
~2.6
Vector
Interface
Control
AND CONTROL SYSTEM (SCS) ......
. . . ......
A-22
. . . . ......
A-24
SM RCS Functional
Description
. . . ......
A-24
CM RCS Functional
Description
. . . ......
~-26
ELECTRICAL POWERSYSTEM . . . . . . . ......
A-28
Introduction
. . . . . . . . . . . . . . . . . .
A-28
iii
.
..I
.
.
I
.,
*“I
-._--
_-.
-
._I_x__~__.
--
---.
-..*---_l----...
--.-.--~
--..
-~
___-___o__ll_.
-__Y._--
--
Part
Page
Functional
Major
Description
Component/Subsystem
Performance
and Design
Operational
Limitations
Systems Test
Meter
Command Module
A2.7
Functional
Data
......
A-32
...........
A-66
and Restrictions
.....
~-67
................
Lighting
A-70
.........
A-72
A-81
...................
A-81
Description
..............
A-83
.................
A-86
TELECOMMUNICATIONS SYSTEM .............
Introduction
Functional
A-88
...................
A-88
Description
..............
A-89
SEQUENTIAL SYSTEMS .................
Introduction
Sequential
Origin
A2.10
Description
Interior
Oxygen Subsystem
~2.9
A-30
ENVIRONMENTAL CONTROL SYSTEM ............
Introduction
A2.8
..............
A-90
...................
Events
A-90
Control
of Signals
Subsystem
A-90
.......
................
~-92
CAUTION AND WARNING SYSTEM .............
Introduction
Functional
Major
...................
A-93
Description
..............
Component Subsystem
Operational
A-93
Limitations
Description
A-93
......
and Restrictions
A-93
.....
A-97
iv
..-l ,....-- _-__---_____-
-,.. -..-
-_
-
-
Page
Part
A2.11
MISCELLANEOUS SYSTEMS DATA . . . . . . . . . . . . .
Introduction
Timers
. . . . . . . . . . . . . . . . . . .
......................
Accelerometer
A-99
A-99
.............
(G-meter)
Command Module
A-?9
Uprighting
System
A-99
. . . . . . . . .
A-99
A2.12
CREWPERSONAL EQUIPMENT ..............
A-103
A2.13
DOCKING AND TRANSFER ................
A-106
Introduction
Functional
A3
. . . . . . . . . . . . . . . . . . .
A-109
LUNAR MODULE SYSTEMS DESCRIPTION ...........
A-111
INTRODUCTION ....................
A-111
LM CONFIGURATION ..................
A-111
Ascent
Descent
Stage
. . . . . . . . . . . . . . . . . . .
A-114
. . . . . . . . . . . . . . . . . .
A-119
Stage
IN - SLA - S-IVB
A4
..............
Description
A-106
A-119
...........
Connections
LM-CSM Interfaces
................
A-119
Stowage Provisions
................
A-122
..........
MISSION CONTROL CENTER ACTIVITIES
A-123
INTRODUCTION ....................
A-123
MISSION OPERATIONS CONTROLROOM ..........
A-128
MCC SUPPORT ROOMS .................
A-131
MISSION SUPPORT AREAS
...............
A-133
Communications,
Command, and Telemetry
System (CCATS) .................
A-133
Y
__~-_-_.-,
.._
. ..-----
.I.,.
.__
,.-__-
-~--_
.-..
--l----l--l-..---
_-“--l_-.
Part
Page
Real-Time
A5
Computer
Complex
(RTCC).
. . . . . . . .
A-134
EXCERPTS FROM APOLLO FUEL CELL AND CRYOGENIC GAS
STORAGE SYSTEM FLIGHT SUPPORT HANDBOOK . . . . . . .
A-139
__.^.
vi
.,___--.-....__I..^
,111-
.._^x_-..-l _..l..-^..l-I_
-__ -.-.--~-
-__~,--
*-_.
---.-
BASELINE DATA:
APOLLO 13 FLIGHT
SYSTEMS AND OPERATIONS
Part Al briefly
describes
Appendix A is divided
into five parts.
the Apollo spacecraft
configuration;
Part A2 provides
a systems description of the Apollo
spacecraft
configuration
with special
emphasis on the
electrical
power system (EPS); Part A3 describes
the lunar module systems;
Part A4 briefly
describes
the Mission
Control
Center at Houston, Texas,
and its interface
with the spacecraft
during the mission;
and Part A5
gives a detailed
description
of the fuel cells
and cryogenic
gas storage
systems aboard the Apollo spacecraft.
This baseline
material
may not
always represent
the precise
Apollo 13 configuration
in every case, since
there is a continuous
updating
which is documented periodically.
For example, Fuel Cell 2 on Apollo
13 was normally
connected to bus A in the
distribution
system, rather
that as described
in Part ~2.6.
The data were extracted
from the
following
sources:
APPENDIX A
PART Al
and A2
Technical
Manual SM2A-03-Block
II-(l)
Apollo Operations
Handbook Block II
Volume 1, dated January 15, 1970.
PART A3
Technical
Handbook,
1970.
Spacecraft,
Manual LMA790-3-L&l, Apollo Operations
Lunar Module, Volume 1, dated February
Center
August
Flight
Operations
31, 1969.
Plan
1,
PART A4
Manned Spacecraft
dated
H Missions,
-
PART A5
Apollo Fuel Cell and Cryogenic
Gas Storage System
Flight
Support Handbook, dated February
18, 1970,
prepared by Propulsion
and Power Division,
Manned
Spacecraft
Center.
A-l
I~..
I .,.
.__ ".----.
-..
. _.---.-l-__-l.~
----
1---
_l.-l--l__
This
page left
blank
intentionally.
.-
A-2
PART Al
APOLLO SPACECRAFT CONFIGURATION
The Apollo
spacecraft
consists
command module (CM), service
module
adapter
(SLA), and the lunar module
stations
are shown in figure
Al-l.
of a launch escape assembly (LEA),
(SM), the spacecraft
lunar module
(LM).
The reference
system and
LAUNCH ESCAPE ASSEMBLY
The LEA (fig.
Al-2) provides
the means for separating
the CM from
the launch vehicle
during pad or first-stage
booster
operation,
This
assembly consists
of a Q-ball
instrumentation
assembly (nose cone),
ballast
compartment,
canard surfaces,
pitch
control
motor, tower jettiskirt,
an open-frame
tower,
son motor, launch escape motor, a structural
and a boost protective
cover (BPC).
The structural
skirt
at the base
which encloses
the launch escape rocket motors,
is
of the housing,
secured to the forward portion
of the tower.
The BPC (fig.
Al-31 is
attached
to the aft end of the tower to protect
the CM from heat during
and from exhaust damage by the launch escape and tower jettison
boost,
secure the tower
Explosive
nuts, one in each tower leg well,
motors.
to the CM structure.
COMIWNDMODULE
The CM (fig.
Al-b),
the spacecraft
control
center,
contains
necessary automatic
and manual equipment to control
and monitor
the spacethe required
equipment for safety
and
craft
systems; it also contains
The module is an irregular-shaped,
primary
comfort of the flight
crew.
structure
encompassed by three heat shields
(coated with ablative
material and joined or fastened
to the primary
structure)
forming
a trunThe CM consists
of a forward
compartment,
a
cated, conic structure.
and an aft compartment for equipment.
(See fig.
Al-b.)
crew compartment,
The command module is conical
shaped, 11 feet 1.5 inches long,
The
12 feet 6.5 inches in diameter without
the ablative
material.
ablative
material
is nonsymmetrical
and adds approximately
4 inches
the height
and 5 inches to the diameter.
A-3
and
to
.
_...
Figure
Al-l.-
Block
II
spacecraft
reference
jr\ - 1-I
.-__.
I. -..
.I --.lr
_..
-- ._ ...--“-_l_*._“.
.-~~.
-.
stations.
Q-MLL
(NOSE
CONE)
PITCH CONTROL
CANARDS
LAUNCH
ESCAPE MOTOR
JETTISON
jj/
--“.
LAUNCH
MOTOR
MOTOR
ESCAPE ASSEMBLY
STRUCTURAL
SKIRT
A..
LAUNCH
TOWER ATTACHMENT
ESCAPE TOWER
COMMAND
(4)
MODULE
BOOST PROTECTIVE
CM-SM
FAIRING
REACTION CONTR
SYSTEM ENGINES
SERVICE MODULE
EC5 RADIATOR
MI
SPS ENGINE
EXPANSION
NOZZLE
SPACECRAFT LM
ADAPTER (SLA) \
NOTE:
LM IS NOT UTILIZED
SOME MISSIONS
Figure
ON
Al-2.-
c
\
Block
II
spacecraft
S-IV8 INSTRUMENT
UNIT
(SHOWN AS REFERENCE)
configuration.
A-5
. .. __.~.
- .,._...- .^_IL-
-.--
..-...-___--_.-
...-...----._l-.-_l”...-l
.._----.~--.“-,_.---_
!:
SlMOd
F4l3N3 MVA
51106 3Nl3N3
1108 538
A. ---~A
5iUOd
3Nl3N3
H31ld
$38
/
IN3A 111 311115
:-
t4oaNl.v
via .8
UOd 3NIW3
HXld
538
COMMAND MODULE
Figure
Al-b.-
Block
II
command module.
SERVICE MODULE
The service
module (fig.
Al-51 is a cylindrical
structure
formed
by l-inch-thick
aluminum honeycomb panels.
Radial beams, from milled
aluminum alloy
plates,
separate
the structure
interior
into six unequal
sectors
around a circular
center section.
Equipment contained
within
-Y
’ccgq
1 and 4 are 50-degree
2 and 5 are 70-degree
3 and 6 are 60-degree
Service
sectors
sectors
sectors
module items
Sector 1
Empty NASA equipment
Sector 2
Environmental system space radiator
Service propulsion system
Reaction control system package (+Y -axis)
Service propulsion system oxidizer sump tank
Sector 4
Fuel cell power plant (three)
Helium servicing panel
Super-critical
oxygen tank (two)
Super-critical
hydrogen tank (two)
Reaction control system control unit
Electrical power system power control relay box
Service module jettison controller sequencer (two)
Sector 5
Environmental control system space radiator
Service propulsion system fuel sump tank
Reaction control system package C-Y axis)
Sector 6
Environmental control system space radiator
Reaction control system package (-Z axis)
Service propulsion system fuel storage tank
Center Section
Sector 3
Service propulsion system helium tank (two)
Service propulsion system
Service propulsion system engine
Reaction control system package (+Z -axis)
Environmental system space radiator
Fairing
Service propulsion system oxidizer storage tank
Electrical power system space radiator’s (eight)
Figure
Al-5.-
Service
id-8
module.
the service
module is accessible
through maintenance
doors located
around
the exterior
surface
of the module.
Specific
items,
such as propulsion
and most of the SC onboard consumables
systems (SPS and RCS), fuel cells,
(and storage tanks)
contained
in the SM compartments,
are listed
in
The service
module is 12 feet 11 inches long (high)
and
figure
Al-5.
12 feet 10 inches in diameter.
Radial beam trusses
on the forward portion
of the SM structure
provide
a means for securing
the CM to the SM. Alternate
beams one,
and five have compression
pads for supporting
the CM. Beams two,
three,
four,
and six have shear-compression
pads and tension
ties.
A flat
center section
in each tension
tie incorporates
redundant
explosive
charges for SM-CM separation.
These beams and separation
devices are
enclosed within
a fairing
(26 inches high and 13 feet in diameter)
between the CM and SM.
SPACECRAFT LM ADAPTER
The spacecraft
LM adapter
(SLA) (fig.
Al-6)
is a large truncated
It houses
cone which connects the CSM and S-IVB on the launch vehicle.
the lunar module (LM), the nozzle of the service
propulsion
system, and
The adapter,
constructed
the high-gain
antenna in the stowed position.
is 154 inches in diameter
at the
of eight 2-inch-thick
aluminum panels,
Separation
forward end (CM interface)
and 260 inches at the aft end.
of the CSM from the SLA is accomplished
by means of explosive
charges
which disengage the four SLA forward panels from the aft portion.
The
individual
panels are restrained
to the aft SLA by hinges and accelerWhen reaching
an
ated in rotation
by pyrotechnic-actuated
thrusters.
angle of 45 degrees measured from the vehicle's
X-axis,
spring thrusters
The panel jettison
velocity
and
(two per panel)
jettison
the panels.
direction
of travel
is such as to minimize
the possibility
of recontact
with the spacecraft
or launch vehicle.
A-9
Panel separation by
explosive charges
FAM-1503F
Figure
Al-6.-
Spacecraft
LM adapter.
PART A2
SYSTEMS DESCRIPTION DATA
INTRODUCTION
Systems description
data include
description
of operations,
component description
and design data, and operational
limitations
and restrictions.
Part 2.1 describes
the overall
spacecraft
navigation,
guidance,
and control
requirements
and the resultant
systems interface.
Parts A2.2
through A2.10 present
data grouped by spacecraft
systems,
arranged in the
following
order:
guidance
and navigation,
stabilization
and control,
service propulsion,
reaction
control,
electrical
power, environmental
control,
telecommunications,
sequential,
and caution
and warnings.
Part A2.11 deals
with miscellaneous
systems data.
Part A2.12 deals with crew personal
equipment.
Part A2.13 deals with docking and crew transfer.
These data were extracted
from the technical
manual SM2A-03BLOCK
(l),
Apollo Operations
Handbook, Block II Spacecraft,
Volume 1, dated
January 15, 1970.
A-11
II-
PART A2.1
GUIDANCE AND CONTROL
Guidance
and Control
Systems
Interface
The Apollo guidance
and control
functions
are performed by the
primary
guidance,
navigation,
and control
system (PGNCS), and stabilization
and control
system (SCS).
The PGNCS and SCS systems contain
rotational
and translational
attitude
and rate sensors which provide
discrete
input information
to control
electronics
which, in turn,
integrate and condition
the information
into control
commands to the spacecraft propulsion
systems.
Spacecraft
attitude
control
is provided
by
commands to the reaction
control
system (RCS).
Major velocity
changes
are provided
by commands to the service
propulsion
system (SPS).
Guidance and control
provides
the following
basic functions:
a.
Attitude
reference
b.
Attitude
control
C.
Thrust
and thrust
vector
control.
-
The basic guidance
and control
functions
may be performed
automatically,
with primary
control
furnished
by the command module computer
(CMC) or manually,
with primary
control
furnished
by the flight
crew.
The subsequent
paragraphs
provide
a general
description
of the basic
functions.
Attitude
Reference
The attitude
reference
function
(fig.
A2.1-1)
provides
display
of
the spacecraft
attitude
with reference
to an established
inertial
refThe display
is provided
by two flight
director
attitude
indierence.
cators
(FDAI) located
on the main display
console,
panels 1 and 2. The
displayed
information
consists
of total
attitude,
attitude
errors,
and
angular
rates.
The total
attitude
is displayed
by the FDAI ball.
Attitude errors
are displayed
by three needles across scales on the top,
right,
and bottom of the apparent periphery
of the ball.
Angular rates
are displayed
by needles across the top right,
and bottom of the FDAI
face.
Total attitude
information
is derived
from the IMU stable platform
or the gyro displ~
coupler
(GDC).
The IMU provides
total
attitude
by
maintaining
a gimbaled,
gyro-stabilized
platform
to an inertial
reference
orientation.
The GDC provides
total
attitude
by updating
attitude
information with angular
rate inputs
from Qyro assembly 1 or 2.
A-12
-__
.___^.-.__-.
-.- _.-..
-_--
._I
I
I INERTIAL
I MEASUREMENT
UNITIIMU)
I _
'
TOTAL ATTITUDE
I
I
GIMBAL
ANGLES
INERTIAL
COUPLING
DATA UNIT
*
(ICDUl
'
COMMAND
MODULE
COMPUTER
(CMC)
ATTITUDE
ERRORS
1
GIMBAL
ANGLES
GDC TOTAL
ATTITUDE
GYRO
DISPLAY
COUPLER
(GDCI
A
.
o-e--PGNCS
ASCP ATTITUDE
ERROR
e
DISPLAY
KEYBOARD
(DSKY)
I
I
I
J
l
I
RATE -1,
ATTITUDE ERROR-1
ATTITUDE ERROR-2
ATTITUDE-1
ATTITUDE-2
ELECTRONIC
DISPLAY
ASSEMBLY
IEDA)
+
CREW INPUTS
1 ATTITUDE ERROR
,
I
4
_----m-----
I--_---L---------I
I
I
I
I
I
I
I
I
I
I
I
I
-,--v--q
-_---L----
ye--------------
w
DIRECTOR
ATTITUDE
INDICATOR
(FDAI) ND. 1
+
-l
ASCP ATTITUDE
ERROR
ATTITUDE
SET
CONTROL
PANEL
(ASCP)
I
I
I
I
t
IMUlGDC
TOTAL
ATTITUDE
r--‘--e-----
I
l
I
I
I
I
' SWITCHING
L _____ I ---I
I
I
DISPLAY ANGULAR RATES
I
I
scs
Figure
A2.1-1.-
Guidance
-
1
and control.
MAIN
DISPLAY
CONSOLE
(MDC)
c
Attitude
error
information
is derived
from three sources,
The
first
source is from zhe II.L? through
the coupling
Jata unit (CDU) which
c-spares
I;?J gimbai angiea with CLIC commanded angles set into the CDU.
Any angular
difference
between the IMU gimbals and the CDU angles is
sent to the FDAI for display
on the attitude
error- needles.
The second
source is from gyro assembly 1 which contains
three (one for each of the
;i, Y, and Z axes) single-degree-of-freedom
attitude
gyros.
Any spacecraft
rotation
about an axis will
offset
the case of a gyro from the
This rotation
is sensed as a displacement
off null,
and a signal
float.
is picked off which is representative
of the magnitude
and direction
of
This signal
is sent to the FDA1 for display
on the attitude
rotation.
error needles.
The third
source is from the GDC which develops attitude
errors
by comparing angular
rate inputs
from gyro assembly 1 or 2 with
an internally
stored orientation.
These data are sent to the FDA1 for
display
on the attitude
error needles.
Angular rates are derived
from either
gyro assembly 1 or 2. Normally,
the no. 2 assembly is used; however, gyro assembly 1 may be
switched
to a backup rate mode if desired.
For developing
rate information,
the gyros are torqued to null when displaced;
thus, they will
produce an output only when the spacecraft
is being rotated.
The output
signals
are sent to the FDA1 for display
on the rate needles and to the
GDC to enable updating
of the spacecraft
attitude.
Attitude
Control
The attitude
control
function
is illustrated
in figure
A2.1-2.
The
control
may be to maintain
a specific
orientation,
or to command small
rotations
or translations.
To maintain
a specific
orientation,
the
attitude
error
signals,
described
in the preceding
paragraph,
are also
r\,uted to the control
reaction
jet on-off
assembly.
These signals
are
conditioned
and applied
to the proper reaction
jet which fires
in the
direction
necessary
to return
the spacecraft
to the desired
attitude.
The attitude
is maintained
within
specified
deadband limits.
The deadband is limited
within
both a rate and attitude
limit
to hold the
spacecraft
excursions
from exceeding
either
an attitude
limit
or angular
the reaction
jets are fired
To maneuver the spacecraft,
rate limit.
automatically
under ccmmand of the CMC or manually
by flight
crew use of
the rotation
control,
control
function
is
In either
case, the attitude
inhibited
until
the maneuver is completed,
Translations
of small magnitude are performed
along the +X axis for fuel settling
of SPS propellants
prior
to burns,
or for a backup deorbit
by manual commands of the translation
control.
An additional
control
is afforded
by enabling
the minimum impulse control
at the lower equipment bay.
The minimum impulse
control
produces one directional
pulse of small magnitude
each time it is
moved from detent.
These small pulses are used to position
the spacecraft
for navigational
sightings.
A-14
-.
ENERGY STORAGE
ENTRY AND
POST LANDING
BATTERY A
POWER GENERATION
’
FUEL CELL
a’I
.
1 I
CRYOGLNtC
SUBSYSTEM
i-
ENTRY AND
POST LANDING
BATTERY B
FUEL CELL
1A
1,mz-l
No-2
I
ENTRY AND
POST LANDING
BATTERY
I
C
4
I
f
PYRO
BATTERY A
PYRO
BATTERY B
I
1
I
1 Oh
\
1 SENSE
-
R/C
CKT
INERATION
I
I
7
POWERCONVERSION
POWER DISTRIBUTION
I
I
I
I
I
I
-’
II
SENSE
CKT
.
j
I
I
I
I
AC
I
I
I
.
INVERTER
NO.
1
I
AC
INVERTER
NO.
3
4
t
I
I
I
+
+,
I
I
I
I
I
I
I
‘II
I I
AC
INVERTER
...
INVERTER
DC&AC
CONTROL
CIRCUITS
INVERTER
PWR
3M4lNB
(RHE5275j
-A70A
POWER DISTRlBUTION
MAIN
BUS
I SWITCH
F-l
(BATA/c)
(MDC-5)
I
Motor switches
51 and 52
CIOIC when main bus tic
witches
54 and 55 ore set to
Bat A/C and Bet B/C.
0
0
I
I
Figure
A2.6l.-
Electrical
MAIN
BUS
TIE SWITCH
(BAT B / C)
55 (MDC-5)
power
I
subsystem
block
2
FCI coon be connected
to
SMbusB6
FC3foSMbusA
DC bus cantlot
circuit
breoksn
ore illurtmted
in
battety
charger
ond CM DC
bus contml
circuits
schematic
diagram.
A-29
_
..~
(
-
.-
I-
--
--
.
Functional
Description
The primary
source of energy is the cryogenic
gas
Energy storage.storage
system that provides
fuel (H2) and oxidizer
(02) to the power
Two hydrogen and two oxygen tanks,
with the assogenerating
system.
are
located
in
the
service
module.
Storage
ciated
controls
and plum.bing,
of reactants
is accomplished
under controlled
cryogenic
temperatures
and
automatic
and manual pressure
control
is provided.
pressures;
Automatic
heating
of the reactants
for repressurization
is dependent on energy demand by the power generating
and/or envircnmental
control
subsystems.
Manual control
can be used when required.
A secondary
source of energy storage
is provided
by five silver
oxide-zinc
batteries
located
in the CM. Three rechargeable
entry and
postlanding
batteries
supply sequencer logic power at all times,
supplemental dc power for peak loads,
all operating
power required
for entry
and postlanding,
and can be connected
to power either
or both pyro cirTwo pyro batteries
provide
energy for activation
of pyro devices
cuits.
throughout
all phases of a mission.
Power generation.Three Bacon-type
fuel cell power plants,
generting power through electrochemical
reaction
of H2 and 02, supply primary
dc power to spacecraft
systems until
CSM separation.
Each power plant
is
capable of normally
supplying
from 400 to 1420 watts at 31 to 27 V dc (at
fuel cell terminals)
to the power distribution
system.
During normal
operation
all three power plants
generate
power, but two are adequate to
complete the mission.
Should two of the three malfunction,
one power
plant
will
insure
successful
mission
termination;
however,
spacecraft
loads must be reduced to operate
within
the limits
of a single
powerplant.
Normal fuel cell connection
to the distribution
system is:
fuel
cell 1 to main dc bus A; fuel cell 2 to main dc busses A and B; ax-&fuel
is provided
for power
cell 3 to main dc bus B. Manual switch control
plant
connection
to the distribution
system, and manual and/or automatic
control
for power plant
isolation
in case of a malfunction.
During the CSM separation
maneuver, the power plants
supply power
through
the SM buses to two SM jettison
control
sequencers.
The sequencers sustain
SM RCS retrofire
during
CSM separation
and fire
the SM
positive
roll
RCS engines 2 seconds after
separation
to stabilize
the SM
during
entry.
Roll engine firing
is terminated
7.5 seconds after
separation.
The power plants
and SM buses are isolated
from the umbilical
The sequencers
are connected
to the SM buses when
through
a SM deadface.
the CM/SM SEP switch
(MDC-2) is activated;
separation
occurs 100 milliseconds after
switch activation.
L
A-30
Primary dc power is converted
into ac by solid
Power conversion.state static
inverters
that provide
115/200-volt
400-cps 3-phase ac power
up to 1250 volt-amperes
each.
The ac power is connected by motor switch
+Uo twc ac buses for distribution
controls
to the ac loads.
One inverter
has the capability
of supplying
all spacecraft
primary
ac power.
One
inverter
can power both buses while the two remaining
inverters
act as
redundant
sources.
However, throughout
the flight,
each bus is powered
by a separate
inverter.
Provisions
are made for inverter
isolation
in
the event of malfunctions.
Inverter
outputs
cannot be phase synchronized;
therefore,
interlocked
motorized
switching
circuits
are incorporated
to prevent
the connection
of two inverters
to the same bus.
three
solid
verter
A second conversion
unit,
the battery
charger,
assures keeping the
entry and postlanding
batteries
in a fully
charged state.
It is a
state device utilizing
dc from the fuel cells
and ac from the into develop charging
voltage,
Power distribution.Distribution
of dc power is accomplished
via
two redundant
dc buses in the service
module which are connected to two
redundant
buses in the command module through a SM deadface,
the CSM
umbilical,
and a CM deadface.
Additional
buses provided
are:
two dc
buses for servicing
nonessential
loads:
a flight
bus for servicing
inflight
telecommunications
equipment;
two battery
buses for distributing
power to sequencers,
gimbal motor controls,
and servicing
the battery
relay bus for power distribution
switching;
and a flight
and postlanding
bus for servicing
some communications
equipment and the postlanding
loads.
via two redundant
ac buses, providing
Three-phase
ac is distributed
bus selection
through switches
in the ac-operated
component circuits.
Fower to the lunar module is provided
through two umbilicals
which
are manually
connected after
completion
of transposition
and docking,
An average of 81 watts dc is provided
to continuous
heaters
in the abort
sensor assembly (ASA), and cycling
heaters
in the landing
radar,
rendezand inertial
measurement unit
(IMU).
Power
vous radar,
S-band antenna,
consumption
with all heaters
operating
simultaneously
is approximately
309 watts.
LM floodlighting
is also powered through the umbilical
for
use during manned lunar module operation
while docked with the CSM.
A dc sensing circuit
monitors
voltage
on each main dc bus, and an
ac sensing circuit
monitors
voltage
on each ac bus.
The dc sensors provide an indication
of an undervoltage
by illuminating
a warning light.
The ac sensors illuminate
a warning light
when high- or low-voltage
limits
are exceeded.
In addition,
the ac sensors activate
an automatic
disconnect of the inverter
from the ac bus during an overvoltage
condition.
conditions
are displayed
by illumination
of an overload
The ac overload
warning light
and are accompanied by a low voltage
light.
Additional
A-31
_-.._
i.._
.- .-r __
_
.. ___.ll-"""
__--___-.- __I--.- 1-w_ _-_-..-_.. --.._,~_-....-F-L-..-s.--_II
----.i. ---1-_.-
1_11_-
sensors monitor
fuel cell overload
and reverse
current
conditions,
viding
an automatic
disconnect,
together
with visual
indications
disconnect
whenever either
condition
is exceeded.
Switches,
meters,
lights,
controlling
and monitoring
all
Major
and talk-back
indicators
functions
of the EPS.
Component/Subsystem
The subsequent
paragraphs
describe
and each of the various
EPS components.
proof the
are provided
for
Description
the
cryogenic
storage
subsystem
Cryogenic
storage.The cryogenic
storage
subsystem (figs.
~2.6-2
and ~2.6-3) supplies
hydrogen to the EPS, and oxygen to the EPS, ECS, and
for initial
LM pressurization.
The two tanks in the hydrogen and oxygen
systems are of sufficient
size to provide
a safe return
from the furthest
point of the mission
on the fluid
remaining
in any one tank.
The physical
data of the cryogenic
storage
subsystem are as follows:
Weight
of usable
cryogenics
(lb/tank)
320 (min)
28 (min)
H2
O2
Design
storage
pressure
bia>
900235
245 (+15,
Minimum
allowable
operating
pressure
(psi4
Approximate
flow rate
at min dq/dm
(+li;5',F
e;viro;m;nt:
lb hran s
150
100
1.71
0.140
-20)
Approximate
quantities
at
minimum heater
--
an~pef~nt~~$ing
(min dq/dm)
45 to 25%
53 to 33%
Initial
pressurization
from fill
to operating
pressures
is accomplished
by GSE. After attaining
operating
pressures,
the cryogenic
fluids
are in a single-phase
condition,
therefore,
completely
homogeneous.
This
avoids sloshing
which could cause sudden pressure
fluctuations,
possible
damage to internal
components,
and prevents
positive
mass quantity
gaugThe single-phase
expulsion
process continues
at nearly
constant
ing.
pressure
and increasing
temperature
above the 2-phase region.
-
A-32
.
+28 VDC
c‘""'"'"
FILL VENTtGSEI (OV-1)
-
;OXYCEN
CRYOPRESS(MDC-2)
02
FILL VALVE (GSE) (OF-l)
1
L
I
07
4
09
-
02 PRESS
IND WC-2
DENSITY
SIGNAL
CONDITIONER
K
DENSITY
SENSOR
PROBE
1
S"R0E FROM:SURGE TANK
PRESSURE
TANK
rWK1
-I I
Q
FAN MOTORS
,
I
L
,
.
CAUTIC
WARN11
SYSTEh
OXYGEN RELIEF VENT
(ORI
,-,
-J
-C
-
I
n
-
"2
/
I
HEATERS t
"2
I
-;(
HEATERS,
'I
1
9
2
CRYOGENIC TAMS
4.
DENSITYSIGNAL
CONDITIONER
,-OXYGEN
I
FILL VALVE IGSEI (OF-Z)
-
07L
I-
~z OXYGEN FILLVENT IGSE) (OV-2)
I,
L
-
1
II
I
II
II
I,
L
-
02
02
Figure
~2.6-2.-
Cryogenic
I
02
02
stcrage
A-34
02
-
0
I
OXYGEN PUPGE VALVE IGSE) (OP)
-
I
I
L
subsystem
(oxygen).
I
I
o2
-
0.
,
& TANK 1 & TANK 2
PRESS XDUCERS
ZRYO PRESS[MDC- 2)
D-
02 TANK NO. 1 MOTOR SWITCH CONTACTS 0
I
I
I
I
I
I
I
I
02 TANK NO. 2
MOTOR SWITCH
CONTACTS
CONTACTS
AL
SM MAIN DC BUS Bl
II
m
p
~
F
’
MAIN D-C
OF
”
t
I
TO. FUEL CELL SHUTOFF VALVES
1
C
I I I I
I I I I
02 VAC ION PUMPS
MN A (RHEB 229)
A
5A”
A
MN 8 IRHEB 229)
05A0
02 HEATERS -2
:ENIC TAMS
CRYOGENIC 02 HTR-2
MNB tRHEB-226)
j
(MDC- 2)
AUTO
( rl’
I L
OFF
ON
PRESSURE AND MOTOR
SWITCHES ARE SHOWN
IN LOW PRESSURE POSITION
0
02
TO
ECS
ON
I
1
.
.
.
I
.
.-._
H? TAN-K 2
DiNSlTY 8 TEhtP
5 IGNAL COND
i
1\
-L
‘\TACTS @
H2TANK 1
DENSITY & TEh4P
SIGNAL COND
0, TANK NO. 2
~TOR sw lTcH
CONTACTS
QTY AMPL 1
AC I (RHER-2261
CRYOGENIC
FAN MOTORS
TANK 1
AC2 (RHEB-2261
02 FANS -1
1MDC-2 I
II
OFF c.+o-
-
A-C BUS
NO. 1
2A -
:RYOGENIC FAN MOTORS
TANK 2
A-C BUS
AC2 (RHEB-226)
ON
-l
I
I
I
H2 TANK 2
DENSITY C TEMP
SIGNAL COND
H2FANS - 2 SWITCH
IMDC - 21
H2FANS - 1 SWITCH
IMDC - 21
d
I
4
HYDROGEN FILL VENT IGSE) IHV-1)
r,
1
HYDROGEN FILL VALVE (GSE) (HF-1)
C
H2
I
I
CONDITIONER
iMPL
I
1
HYDROGEN RELIEF! VENT (FLYAWAY UMBILICAL)
\
-
H2
/
(HR)
H2
CRYOGENIC TANKS
c
ION
PUMP
HYDROGEN FILL VALVE (CSE) (HF-21
r\
I,
Li
-
H2
H2
H2
HYDROGEN FILL VENT IGSE) IHV-2)
-
I
HYDROGEN PURGE VALVE IGSE) (HP)
H2
-
H2
FROM CB
CRYOGENIC FAN MOTORS TANK 2 AC2 (RHEE-226)
CRYOGENIC FAN MOTORS TANK 1 AC1 (RHEB-2261
9
H2
‘OGEN
TEMP
SENSOR
TEMP
SIGNAL
CONOITIONFR
I
ITT”
SM MAIN DC
I I
-I
I-
CM r)-CMAlN
lVAC PS/,.q
ION
KiNAD”
PUMP
I;;y=
II
VALVE
A
-
H2
Hz -
q$F//glm
._ ._.
[
TANKsm
IRE
I-
D-&IN
BUS B
,
0
TANK NO. 2
PRESSURE
SWITCH
1
FILTER
Hz
-
I
H2
H2
-
H2
H2
CRYOGENIC H2 HTR 2 MN B (RHEB-226)
Ii2 TANK NO. 1 MOTOR tiITCH
CONTACTS (j)
Ii2 TANK NO. 2
MOTOR SWITCH
, CONTACTS
_
1
H FANS -1
L DC-2)
AUTO
6C
OFF v
/
-C MAIN
BUS A
\
CRYOGENIC H2 HTR lMN A IRHEB-226)
H2 HEATERS - 1
(MDC-2)
AUTO
\
w OFF
0
- H2 HEATERS -
\
CRYOGENIC H2 HlR 2MN B (RHEB-22%)
-
H2 FANS --,
H2 FANS -2
(MDC-2)
AUTO
*A
OFF wT+
71
---A
0
OFF
-1
ON
---
01
02
Figure
PRESSUREAND MOTOR SWITCHES ARE
SHOWN IN LOW PRESSUREPOSITION
VAC ION PUMP FUSES OPENED
DURING PRELAUNCH COUNTDOWN
~2.6-3.-
Cryogenic
storage
A-35
subsystem
(hydrogen)
Two parallel
dc heaters
in each tank supply the heat necessary
to
maintain
design pressures.
Two parallel
3-phase ac circulating
fans
circulate
the fluid
over the heating
elements to maintain
a uniform
density and decrease the probability
of stratification.
A typical
heater
and fan installation
is shown in figure
A2.6-4.
Relief
valves provide
check_valves
provide
tank isolation,
and individual
overpressure
relief,
fuel cell shutoff
valves provide
isolation
of malfunctioning
power plants.
Filters
extract
particles
from the flowing
fluid
to protect
the ECS and
EPS components.
The pressure
transducers
and temperature
probes indicate
the thermodynamic
state of the fluid.
A capacitive
quantity
probe indicates
quantity
of fluid
remaining
in the tanks.
FAN &MOTOR
CAPYl1lvE
PRORE-
‘I
v
ENCASED
INTERNALLY
Figure
A2.6-4.-
Cryogenic
pressurization
measurement devices.
and quantity
c
,
A-36
or manually
Repressurization
of the systems can be automatically
The automatic
mode is designed to give
controlled
by switch selection.
at design pressures.
a single-phase
reactant
flow into the feed lines
controlled
through a pressure
The heaters
and fans are automatically
As pressure
in the tanks decreases,
switch-motor
switch arrangement.
the pressure
switch in each tank closes to energize
the motor switch,
Both tanks have to declosing
contacts
in the heater and fan circuits.
When
crease in pressure
before heater and fan circuits
are energized.
either
tank reaches the upper operating
pressure
limit,
that respective
pressure
switch opens to again energize
the motor switch,
thus opening
The 02 circuits
are energized
the heater and fan circuits
to both tanks.
at
865 psia minimum and de-energized
at 935 psia
maximum.
The H2 circuits
The
energize
at 225 psia minimum and de-energize
at 260 psia maximum.
most accurate
quantity
readout will
be acquired
shortly
after
the fans
During all other periods
partial
stratification
may dehave stopped.
grade quantity
readout accuracy.
When the systems reach the point where heater and fan cycling
is at
a minimum (due to a reduced heat requirement),
heat leak of the tank is
sufficient
to maintain
design pressures,
provided
flow is within
the min
This realm of operation
dq/dm values shown in the preceding
tabulation.
The minimum heat requirement
is referred
to as the min dq/dm region.
region
for oxygen starts
at approximately
45-percent
quantity
and terBetween these tank quanminates at approximately
25-percent
quantity.
tities,
minimum heater and fan cycling
will
occur under normal usage.
The amount of heat required
for repressurization
at quantities
below
25-percent
starts
to increase
until
below the j-percent
level
practically
In the hydrogen system,
continuous
heater and fan operation
is required.
the quantity
levels
for minimum heater and fan cycling
are between apwith continuous
operation
occurring
at
53 and 33 percent,
proximately
approximately
the 5 percent-level.
Assuming
a constant
level
flow
from each tank
(02 - 1 lb/hr,
0.09 lb/hr)
each successive
repressurization
period
is of longer
H2 The periods
between repressurizations
lengthen
as quantity
duration.
decreases
from full
to the minimum dq/dm level,
and become shorter
as
quantity
decreases from the minimum dq/dm level
to the residual
level.
Approximate
repressurization
periods
are shown in table A2.6-I,
which
also shows the maximum flow rate in pounds per hour from a single
tank
with the repressurization
circuits
maintaining
minimum design pressure.
The maximum continuous
flow that each cryogenic
tank can provide
at
minimum design pressure
is dependent on the quantity
level and the heat
The heat required
to maintain
a conrequired
to maintain
that pressure.
stant pressure
decreases as quantity
decreases
from full
to the minimum
A-37
As quantity
decreases beyond the minimum dq/dm region,
the
dq/dm point.
heat required
to maintain
a constant
pressure
increases.
As fluid
is
a specific
amount of heat is withdrawn.
withdrawn,
When the withdrawal
rate exceeds the heat that can be supplied
by the heaters,
fan motors,
and heat leak,
there is a resultant
pressure
decrease below the minimum
design operating
level.
The ability
to sustain
pressure
and flow is a factor
of heat required
versus the heat provided
by heaters,
fan
Since heat leak characteristics
of each tank
heat leak.
the flow each tank can provide
will
also vary to a small
input from heaters,
fan motors,
and heat leak into an 02
595.87
motors
similar
of the amount
motors,
and
vary slightly,
degree.
Heat
tank is
Btu/hour
(113.88-watt
heaters
supply 389.67 Btu, 52.8-watt
fan
supply 180.2 Btu, and heat leak supplies
26 Btu).
Heat input from
(18.6-watt
heaters
supply
sources into a H2 tank is 94.6 Btu/hr
fan motors supply 23.89 Btu, and heat leak supplies
63.48 Btu, 7-watt
These figures
take into consideration
the line loss between
7.24 Btu).
the power source and the operating
component.
TABLE ~2.6-I.Quantity
(percent)
100
:z
85
80
75
:p
60
55
z;
40
52
25
20
1.5
10
'I . 5
5
OXYGENAND HYDROGENREPRESSURIZATION AND FLOW.
Wren
Repressurization
time, minutes
(865 to 935 psia)
4.0
4.3
4.6
Flow at
865 psia
3.56
I
5::
2:;
7.4
8.7
9.6
10.8
11.5
12.4
12.6
13.0
13.1
'4.;:
.
5.27
6.02
7.01
7.94
9.01
10.80
12.54
14.19
15.69
17.01
17.56
17.56
16.55
20.0
21.0
22.0
23.0
24.5
26.5
28.5
31.0
33.5
36.0
39.0
41.0
41.0
41.0
40.5
40.5
42.0
47.0
58.0
71.0
Continuous
0.38
0.42
0.46
0.49
0.52
0.65
0.76
0.80
0.87
0.93
0.97
0.98
0.97
0.94
0.91
0.83
0.71
0.54
0.37
0.23
0.16
A-38
-.
--..
I..-.--
-I_i_^-----_-.--~ll~-IIII1l~.
__;__-.
__.I.~-
--.-
.I____
I__
-
To avoid excessive
temperatures,
which could be realized
during
continuous
heater and fan operation
at extremely
low quantity
levels,
a
thermal
sensitive
interlock
device is in series
with each heater element.
The device automatically
opens the heater circuits
when internal
tank
at +7G" F.
shell
temperatures
reach +9G" F., and closes the circuits
oxygen temperature
will
be approximately
Assuming normal consumption,
-157" F., at mission
termination,
while hydrogen temperature
will
be
approximately
-385" F.
The manual mode of operation
bypasses the pressure
switches,
and
supplies
power directly
to the heaters
and/or fans through the individual
control
switches.
It can be used in case of automatic
control
failure,
heater failure,
or fan failure.
Tank pressures
and quantities
are monitored
on meters located
on
The caution
and warning system (CRY0 PRESS) will
alarm when
MIX-2.
oxygen pressure
in either
tank exceeds 950 psia or falls
below 8GG psia.
Since a
The hydrogen system alarms above 270 psia and below 220 psia.
common lamp is provided,
reference
must be made to the individual
pressure
and quantity
meters (MDC-2) to determine
the malfunctioning
tank.
Tank
and
reactant
temperatures
of
each
tank
are
telempressures, quantities,
etered to MSFN.
Oxygen relief
valves vent at a pressure
between 983 and 1010 psig
and reseat at 965 psig minimum.
Hydrogen relief
valves vent at a pressure
Full flow
between 273 and 285 psig, and reseat at 268 psig minimum.
venting
occurs approximately
2 pounds above relief
valve opening pressure.
All the reactant
tanks have vat-ion
pumps to maintain
the integrity
of the vacuum between the inner and outer shell,
thus maintaining
heat
SM main de bus A distributes
power
leak at or below the design level.
to the H2 tank 1 pump and bus B to the H2 tank 2 pump. Fuses provide
power source protection.
disable
the circuit
for
MNA- MJTB (m-29),
buses, which distribute
breakers
allow
These fuses are removed during prelaunch
to
Circuit
breakers,
G2 VAC ION PUMPS flight.
for the CM main
Provide power source protection
The circuit
power to the O2 vat-ion
pumps.
use of the O2 vat-ion
provide a means of disabling
are opened on the launch
pad,
circuit
pump circuits
if necessary.
and closed
throughout
flight,
The O2 circuit
at 90 percent
tank
and
breakers
quantity.
The most likely
period of overpressurization
in the cryogenic
system
The possibility
will
occur during operation
in the minimum dq/dm region.
of overpressurization
is predicted
on the assumption
of a vacuum breakAlso, under certain
condown, resulting
in an increase
in heat leak.
extremely
low
power
levels
and/or
a
depressurized
cabin,
ditions,
that is,
A-39
~*-_....
-- __ ‘.__”-.,.. _.-.--.
--. . _...,“.---.
__cI_ -- .-__..
--.x_I-
demand may be lower
preceding
conditions
tank.
than the minimum dq/dm flow
would result
in an increase
necessary.
Any of the
of pressure
within
a
In the case of hydrogen tank overpressurization,
prior
to reaching
relief
valve cracking
pressure,
tank pressure
can be decreased by performing
an unscheduled
fuel cell hydrogen purge.
A second method for
relieving
overpressure
is to increase
electrical
loads,
thus increasing
fuel cell demand.
However, in using this method, consideration
must be
given to the fact that there will
be an increase
in oxygen consumption,
which may not be desirable.
Several procedures
can be used to correct
an overpressure
condition
in the oxygen system.
One is to perform an unscheduled
fuel cell purge.
A second is to increase
oxygen flow into the command module by opening
the ECS DIRECT O2 valve.
The third
is to increase
electrical
loads,
which may not be desirable
gen consumption.
because
this
method will
A requirement
for an overpressure
correction
tems simultaneously
is remote, since both reactant
the minimum dq/dm region in parallel.
also
increase
in both reactant
systems do not
hydrosysreach
.
During all missions,
to retain
a single
tank return
capability,
there
is a requirement
to maintain
a balance between the two tanks in each of
the reactant
systems.
When a 2- to b-percent
difference
is indicated
on
the oxygen quantity
meters (MIX-2),
the O2 HEATERS switch (MDC-2) of the
lesser
tank is positioned
to OFF until
tank quantities
equalize.
A
3-percent
difference
in the hydrogen quantity
meters (MDC-2) will
require
positioning
the H2 BEATERS switch
(MDC-2) of the lesser
tank to OFF until
tank quantities
equalize.
This procedure
retains
the automatic
operation
of the repressurization
circuits,
and provides
for operation
of the fan
motors during repressurization
to retain
an accurate
quantity
readout in
all tanks.
The necessity
for balancing
should be determined
shortly
after
a repressurization
cycle,
since quantity
readouts
will be most accurate
at
this time.
Batteries.Five silver
in the EPS. These batteries
oxide-zinc
are located
storage batteries
in the CM lower
are incorporated
equipment bay.
Three rechargeable
entry and postlanding
batteries
(A, B, and C)
power the CM systems after
CSM separation
and during postlanding.
Prior
to CBI separation,
the batteries
provide
a secondary source of power while
The entry batteries
are used for
the fuel cells
are the primary
source.
the following
purposes:
A-40
a.
Provide
CM power after
Supplement
b.
maneuvers)
fuel
cell
CSM separation
power during
C.
Provide
power during
emergency
d.
Provide
power for
EPS control
e.
Provide
sequencer
logic
f.
Provide
power for
recovery
g*
Batteries
peak load
operations
periods
(failure
(Delta
V
of two fuel
cells)
circuitry
(relays,
indicators,
etc.)
power
aids
during
A, B, or C can power pyro
postlanding
circuits
by selection.
Each entr;{ and postlanding
battery
is mounted in a vented plastic
case and consists
of 20 silver
oxide-zinc
cells
connected in series.
The cells
are individually
encased in plastic
containers
which contain
relief
valves that open at 35 + 5 psig, venting
during an overpressure
into the battery
case.
The three cases can be vented overboard
through
a common manifold,
the BATTERY VENT valve (RHEB-252), and the ECS waste
water dump line.
Since the BATTERY VENT is closed prior
to lift-off,
the interior
of
the battery
cases is at a pressure
of one atmosphere.
The pressure
is
relieved
after
earth orbit
insertion
and completion
of cabin purge by
After
completion
the
positioning
the control
to VENT for 5 seconds.
control
is closed,
and pressure
as read out on position
4A of the System
Test Meter (LEB-101) should remain at zero unless there is battery
outGutgassing
can be caused by an internal
battery
failure,
an
gassing.
abnormal high-rate
discharge,
or by overcharging.
If a pressure
increase
is noted on the system test meter, the BATTERY VENT is positioned
to VENT
for 5 seconds, and reclosed.
Normal battery
charging
procedures
require
a check of the battery
manifold
after
completion
of each recharge.
Since
the battery
vent
line
is
connected
line,
it provides a means of monitoring
which
would
be
indicated
by
when the BATTERY VENT control
to the waste
water
dump
waste water dump line plugging,
a pressure
rise
is positioned
in the battery
to VENT.
manifold
line
Each battery
is rated at 40-ampere hours (AH) minimum and will
deliver
this at a current
output of 35 amps for 30 minutes and a subsequent
output of 2 amps for the remainder
of the rating.
A-41
At Apollo mission
and will
provide
this
ever, 40 AH is used in
45 AH for postlanding
loads,
each battery
is capable of providing
amount after
each complete recharge
cycle.
mission
planning
for inflight
capability,
capability
of a fully
charged battery.
45 AH
Howand
Open circuit
voltage
is 37.2 volts.
Sustained
battery
loads are
a battery
bus voltage
of
extremely
light
(2 to 3 watts);
therefore,
approximately
34 V dc will
be indicated
on the spacecraft
voltmeter,
exhave been activated
to tie the battery
cept when the main bus tie switches
only batteries
A and B will
be
outputs
to the main dc buses.
Normally,
Battery
C is isolated
during prelaunch
connected
to the main de buses.
by opening the MAIN A-BAT C and MAIN B-BAT C circuit
breakers
(RHEB-275).
Battery
C will
therefore
provide
a backup for main dc bus power in case
of failure
of battery
A or B or during
the time battery
A or B is being
The two-battery
configuration
provides
more efficient
use of
recharged.
fuel cell power during peak power loads and decreases overall
battery
The M&IN A- and MAIN B-BAT C circuit
breakers
are closed
recharge
time.
prior
to CSM separation
or as required
during recharge
of battery
A or B.
Battery
C, through
circuit
breakers
BAT C to
BAT BUS B (RHEB-250), provides
backup power to the
in the event of failure
of entry battery
A or B.
are normally
open until
a failure
of battery
A or
can also be used to recharge battery
A or B in the
the normal charging
circuit.
BAT BUS A and BAT C to
respective
battery
bus
These circuit
breakers
This circuit
B occurs.
event of a failure
in
-
The two pyrotechnic
batteries
supply power to initiate
ordnance debatteries
are isolated
from the rest of
vices in the SC. The pyrotechnic
the EPS to prevent
the high-power
surges in the pyrotechnic
system from
These
affecting
the EPS, and to insure
source power when required.
Entry and postlanding
batteries
are not to be recharged
in flight.
oattery
A, B, or C can be used as a redundant
source of power for initiating
pyro circuits
in the respective
A or B pyro system, if either
This can be performed by proper manipulation
of the
pyro battery
fails,
Caution must be exercised
to isolate
the
circuit
breakers
on RHEB-250.
failed
pyro battery
by opening the PYRO A (B) SE& A (B) circuit
breaker,
prior
to closing
the yellow
colored BAT BUS A (B) to PYRO BUS TIE circuit
breaker.
-
A-42
Performance
Battery
characteristics
Rated
capacity
per
battery
of each SC battery
open
circuit
voltage
(max.)
Nominal
voltage
are as follows:
Minimum
voltage
Ambient
battery
temperature
Entry and
Postlanding
A, B, and
c (3)
40 amp-hrs
(25 ampere
rate)
37.8 V dc mar 29 V dc 27 V dc
(37.2 V dc i; (35 amps (35 amps
50” to
110" F
Pyre A and
B (2)
0.75 amphrs (75
amps for
36 seconds)
37.8 V dc max 23 v dc 20 v dc
(37.2 V dc ir (75 amps (75 amps
60” to
110" F
flight)
flight)
load)
load)
NOTE: Pyro battery
load voltage
is not measurable
extremely
short time they power pyro loads.
load)
load)
(32 V dc
open
circuit)
in the SC due to the
Fuel cell power plants.Each of the three Bacon-type
fuel cell
power plants
is individually
coupled to a heat rejection
(radiator)
system, the hydrogen and oxygen cryogenic
storage
systems, a water storage
system, and a power distribution
system.
A typical
power plant
schematic
A2.L5.
is shown in figure
The power plants
generate
dc power
chemical reaction.
The by-product
water
age tank in the CM where it is used for
cooling
p'zposes
in the ECS. The amount
to the power produced which is relative
table ~2.6-11.)
on demand through an exothermic
is fed to a potable
water storastronaut
consumption
and for
of water produced is equivalent
to the reactant
consumed.
(See
TABLE ~2.6-II.Load
(amps 1
REACTANTCONSWTIOPJurn WATERPRODUCTION
O2 lb/hr
0.0102
T
H2 lb/hr
H2°
Ib/hr
0.001285
0.01149
T
0.0204
0.002570
0.02297
10.42
2
0.0408
0.005140
0.04594
20.84
3
4
0.0612
0.007710
0.0689i
31.26
0.0816
0.010280
0.09188
41.68
5
6
0.1020
0.012850
0.11485
52.10
0.1224
0.015420
0.13782
62.52
7
0.1428
0.017990
0.16079
72.94
8
0.1632
0.020560
0.18376
83.36
9
10
0.1836
0.023130
0.20673
0.2040
0.025700
0.2297
93.78
104.20
15
20
0.3060
0.038550
0.34455
156.30
0.4080
0.051400
0.45940
208.40
25
0.5100
0.064250
260.50
30
0.6120
0.077100
0.57425
0.68910
35
40
0.7140
0.089950
0.8160
0.10280
'15
0.9180
0.11565
50
1.0200
0.12850
55
60
1.1220
0.14135
1.2240
0.15420
65
1.3260
0.16705
1.49305
70
i .4280
0.17990
1.6079
677.30
729.40
75
1.5300
0.19275
1.72275
781.50
80
1.6320
0.20560
1.83760
833.60
85
1.7340
0.21845
885.70
90
1.8360
0.23130
1.95245
2.06730
95
100
1.9380
0.24415
2.18215
2.0400
0.25700
2.2970
O2 = 2.04 x lO-2
I
H = 2.57 x 10 -3 I
2
H20 = 10.42
---~~--___-
3~2.60
0.80395
0.91880
364.70
I.03365
1.1485
448.90
1.26335
1.3782
573.10
cc/amp/hr
H20 = 2.297 x 10m2 lb/amp/hr
A-44
.._.
5.21
0.5
1
XMIJLAS:
-----.“.~
x/hr
416.80
521.00
625.20
937.90
989.00
1042.00
I
1111II
IIIII
1111II
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
uxYGEN GAS PURITY LEVEL (% BY VOLUME)
1,i e;ure AZ. 6- '(.-
O2 gas purity
A-49
effect
on purge
interval.
tiiii
i i
i
I
I
/
/ //
///:
/‘A%’
I
/Ml
I I
I
SPiC I-klRIfY
HYDROGEN GAS INERT LEVEL (P;hl)
II I
10,000
0.1 11 1
99.0
libb
100
99.9
99.99
10
I
99.999
HYDROGEN GAS PURITY LEVEL f% BY VOLUME)
Figure
A2.6-8.-
H2 gas purity
effect
on purge
interval.
-
A-50
I
- . ..-.... 1 . __l_"-^"--.-
.---..-_.
resulting
from a decreased
voltage,
crease in fuel cell skin temperature
voltage.
by a gradual
deload, is followed
which causes a decrease in terminal
The range in which the terminal
voltage
is permitted
to vary is determined by the high and low voltage
input design limits
of the compoFor most components the limits
are 30 volts
dc and
nents being powered.
To remain within
these design limits,
the dc bus voltage
25 volts
de.
must be maintained
between 31.0 and 26.2 volts
dc.
To compensate for
it is recommended sustained
bus voltage
be maintained
becyclic
loads,
Bus voltage
is maintained
within
prescribed
tween 26.5 and 30.0 V dc.
limits
by the application
of entry and postlanding
batteries
during load
increases
(power up).
Load increase
or decrease falls
well within
the
limits
of power supply capability
and, under normal conditions,
should
not require
other than normal checklist
procedures.
Power up.- Powering Up SpaCeCraft systems is performed
in one continuous
sequence providing
the main bus voltage
does not decrease below
26.5 volts.
If bus voltage
decreases to this level,
the power up sequence
can be interrupted
for the time required
for fuel cell temperatures
to
increase
with the resultant
voltage
increase
or the batteries
can be connected to the main buses thus reducing
the fuel cell load.
In most cases,
powering up can be performed
in one continuous
sequence; however, when
starting
from an extremely
low spacecraft
load, it is probable
that a
power up interruption
or earlier
battery
coupling
may be required.
The
greatest
load increase
occurs while powering up for a delta V maneuver.
Power down.- Powering down spacecraft
systems is performed
in one
continuous
sequence providing
the main bus voltage
does not increase
Powering down from relatively
high spacecraft
load
above 31.0 volts.
levels,
that is, following
a delta V, the sequence may have to be interrupted for the time required
for fuel cell temperature,
and as a result,
bus voltage
to decrease.
To expedite
power down, one fuel cell can be
disconnected
from the buses increasing
the loads on the remaining
fuel
cells and decreasing
bus voltage,
thus allowing
continuation
of the power
down sequence.
Fuel cell disconnect.If the requirement
arises to maintain
a
power plant on open circuit,
temperature
decay would occur at an average
6 deg/hr, with the automatic
in-line
heater cirrate of approximately
power
cuit activating
at a skin temperature
of 385’ F and maintaining
plant
temperature
at 385” F. In-line
heater activation
can be confirmed
by a 4.5- to 6-amp indication
as observed on the dc amps meter (MDC-3)
with the de indicator
switch positioned
to the open circuited
fuel cell
position.
Reactant valves remain open.
Fuel cell pumps can be turned
off until
the in-line
heater circuit
activates,
at which time they must
be on.
A-51
Closing
of reactant
valves during a power plant disconnect
is dependent on the failure
experienced.
If power plant failure
is such as
to allow future
use, that is, shutdown due to partially
degraded output,
it is recommended the reactant
valves remain open to provide
a positive
reactant
pressure.
The valves should be closed after
power-plant
skin
temperature
decays below 300" F. The reactant
valves are closed during
initial
shutdown,
if the failure
is a reactant
leak, an abnormally
high
regulator
output pressure,
or complete power-plant
failure.
Prior
to disconnecting
a fuel cell,
if a single
inverter
is being
power plants
is connected to both main de
used, each of the remaining
buses to enhance load sharing
since bus loads are unbalanced.
If two
inverters
are being used, main de bus loads are relatively
equal;
therefore,
each of the remaining
power plants
is connected to a separate main
dc bus for bus isolation.
If one power plant had been placed on open
circuit
for an extended period of time , prior
to powering up to a configuration
requiring
three power plants,
reconnecting
is accomplished
prior
to the time of heavy load demands.
This permits proper conditioning
of the power plant which has been on open circuit.
for
The time required
proper conditioning
is a function
of skin temperature
increase
and the
load applied
to the power plant.
Inverters.Each inverter
(fig.
~2.6-9) is composed of an oscillator,
an eight-stage
digital
countdown section,
a de line filter,
two silicon
controlled
rectifiers,
a magnetic amplifier,
a buck-boost
amplifier,
a
demodulator,
two dc filters,
an eight-stage
power inversion
section,
a
harmonic neutralization
transformer,
an ac output filter,
current
sensing
transformers,
a Zener diode reference
bridge,
a low-voltage
control,
and
an overcurrent
trip
circuit.
The inverter
normally
uses a 6.bkHz square
wave synchronizing
signal
from the central
timing equipment
(CTE) which
maintains
inverter
output at 400 Hz.
If this external
signal
is completely
lost,
the free running
oscillator
within
the inverter
will
provide
pulses that will maintain
inverter
output within
f7 Hz. The internal
oscillator
is normally
synchronized
by the external
pulse.
The subsequent
paragraphs
describe
the function
of the various
stages of the inverter.
._
The 6.4-kHz square wave provided
by the CTE is applied
through the
internal
oscillator
to the eight-stage
digital
countdown section.
The
oscillator
has two divider
circuits
which provide
a 1600-Hz signal
to the
magnetic amplifier.
The eight-stage
digital
countdown section,
triggered
by the 6.4-kHz
displaced
one
signal,
produces eight 400-Hz square waves, each mutually
One pulse-time
is
pulse-time
from the preceding
and following
wave.
The eight square
156 microseconds
and represents
22.5 electrical
degrees.
waves are applied
to the eight-stage
power inversion
section.
..-
25-30 VOCTS
D-C INPUT
I
r - I---
------
7 BlAS
r
------
1
TRANSFORMER
VOLTAGE
A-C
FILTER
DEMODULATOR
VOLTAGE'&
CURRENT
REGULATION
I .6 KH,
r ---
30
I
--s---
OSCILLATOR
CONTROL
I
L-L
BUS 2
-I
I
1
I
I
I
I
I
I
COUNTDOWN
WI
EMP HI4-
:aw
4.4 KHz
rIf--
NEGATIVE
SQUARE WAVE
Figure
~2.6-9. - Inverter
A-53
I--
-------
NOTE:
I
I
I
-- -I
Unless othcwiw
specified:
I. hetier
1 is shown.
2. A denotes input voltage.
block
diagram.
1
The eight-stage
power inversion
section,
fed a controlled
voltage
from the buck-boost
amplifier,
amplifies
the eight 400-Hz square waves
produced by the eight-stage
digital
countdown section.
The amplified
mutually
displaced
22.5 electrical
degrees,
are next
square waves, still
applied
to the harmonic neutralization
transformer.
The harmonic neutralization
section
consists
of 31 transformer
windings
on one core.
This section
accepts the 400-Hz square-wave
output
of the eight-stage
power inversion
section
and transforms
it into a
?-phase 400-Hz 115-volt
signal.
The manner in which these transformers
are wound on a single
core produces flux cancellation
which eliminates
all harmonics
up to and including
the fifteenth
of the fundamental
frequency.
The 22.5" displacement
of the square waves provides
a means of
electrically
rotating
the square wave excited
primary
windings
around the
3-phase,
wye-connected
secondary
windings,
thus producing
the 3-phase
400-Hz sine wave output.
This 115-volt
signal
is then applied
to the ac
output
filter.
The ac output filter
eliminates
the remaining
higher harmonics.
Since the lower harmonics
were eliminated
by the harmonic neutral
transCircuitry
former,
the size and weight of this output filter
was reduced.
in this filter
also produces a rectified
signal
which is applied
to the
The amplitude
of
Zener diode reference
bridge
for voltage
regulation.
After
this signal
is a function
of the amplitude
of ac output voltage.
filtering,
the 3-phase 115-volt
ac 400-Hz sine wave is applied
to the ac
buses through
individual
phase current-sensing
transformers.
The current-sensing
transformers
produce a rectified
signal,
the
amplitude
of which is a direct
function
of inverter
output
current
magnitude.
This dc signal
is applied
to the Zener diode reference
bridge
to
it is also paralleled
to an overcurrent
regulate
inverter
current
output:
sensing circuit.
The Zener diode reference
bridge receives
a rectified
de signal,
from the circuitry
in the ac output filter.
representing
voltage
output,
A variance
in voltage
output unbalances
the bridge,
providing
an error
signal
of proper polarity
and magnitude
to the buck-boost
amplifier
via
the magnetic amplifier.
The buck-boost
amplifier,
through
its bias voltcompensates for voltage
variations.
When inverter
current
age output,
output reaches 200 to 250 percent
of rated current,
the rectified
signal
applied
to the bridge
from the current
sensing transformers
is of sufficient
magnitude
to provide
an error
signaL,causing
the buck-boost
amplifier
to operate in the same manner as during an overvoltage
condition.
The bias output of the buck-boost
amplifier,
controlled
by the error
signal, will be varied
to correct
for any variation
in inverter
voltage
or a
When inverter
current
output
beyond-tolerance
increase
in current
output.
the overcurrent
sensing circuit
is
exceeds 250 percent
of rated current,
activated.
A-54
------__1_1-^-".---. ___..---
-_.
-I_
_--
The overcurrent
sensing circuit
monitors
a rectified
dc signal
repWhen total
inverter
current
output exceeds
resenting
current
output.
this circuit
will
illuminate
an overload
250 percent of rated current,
If
current
output
of
any
single
phase
exceeds
lamp in 15+5 seconds.
300 percent
of rated current,
this circuit
will
illuminate
the overload
lamp in 5-+1 seconds.
The AC BUS 1 OVERLOADand AC BUS 2 OVERLOADlamps
are in the caution/warning
matrix
on MDC-2.
The de power to the inverter
is supplied
from the main dc buses
The filter
reduces the high-frequency
ripple
through
the de line filter.
and the 25 to 30 volts
de is applied
to two siliconin the input,
controlled
rectifiers.
The silicon-controlled
rectifiers
are alternately
set by the 1600-Hz
signal
from the magnetic amplifier
to produce a de square wave with an
on-time
of greater
than 90" from each rectifier.
This is filtered
and
supplied
to the buck-boost
amplifier
where it is transformer-coupled
with
the amplified
1600-Hz output of the magnetic amplifier,
to develop a fildc which is used for amplification
in the power inversion
tered 35 volts
stages.
The buck-boost
amplifier
also provides
a variable
bias voltage
to
the eight-stage
power inversion
section.
The amplitude
of this bias
voltage
is controlled
by the amplitude
and polarity
of the feedback signal from the Zener diode reference
bridge which is referenced
to output
voltage
and current.
This bias signal
is varied
by the error
signal
to
regulate
inverter
voltage
and maintain
current
output within
tolerance.
The demodulator
circuit
compensates for any low-frequency
ripple
(10 to 1000 Hz) in the dc input to the inverter.
The high-frequency
ripple
is attenuated
by the input filters.
The demodulator
senses the
35-volt
dc output of the buck-boost
amplifier
and the current
input to
the buck-boost
amplifier.
An input dc voltage
drop or increase
will
be
reflected
in a drop or increase
in the 35-volt
dc output of the buckboost amplifier,
as well as a drop or increase
in current
input to the
buck-boost
amplifier.
A sensed decrease in the buck-boost
amplifier
voltage
output is compensated for by a demodulator
output,
coupled through
the magnetic amplifier
to the silicon-controlled
rectifiers.
The demodulator
output causes the SCR's to conduct for a longer time, thus increasing
their
filtered
dc output.
A sensed increase
in buck-boost
amplifier
voltage
output,
caused by an increase
in dc input to the inverter,
is compensated for by a demodulator
output coupled through the magnetic
amplifier
to the silicon-controlled
rectifiers,
causing them to conduct
for shorter
periods,
thus producing
a lower filtered
dc output
to the
buck-boost
amplifier.
In this manner, the 35-volt
dc input to the power
inversion
section
is maintained
at a relatively
constant
level
irrespective of the fluctuations
in de input voltage
to the inverter.
A-55
The low-voltage
control
circuit
samples the input voltage
to the
Since the buck-boost
inverter
and can terminate
inverter
operation.
amplifier
provides
a boost action
during a decrease in input voltage
to
35 volts dc to the
in an attempt
to maintain
a constant
the inverter,
power inversion
section
and a regulated
115-volt
inverter
output,
the
high boost required
during a low-voltage
input would tend to overheat
As a precautionary
measure, the
the solid
state buck-boost
amplifier.
low-voltage
control
will
terminate
inverter
operation
by disconnecting
operating
voltage
to the magnetic amplifier
and the first
power inversion
dc.
stage when input voltage
decreases
to between 16 and 19 volts
A temperature
sensor with a range
in each inverter
and provides
an input
a light
at an inverter
overtemperature
is telemetered
to MSFN.
of +32O to +248' F is installed
to the C&WSwhich will
illuminate
of 190" F. Inverter
temperature
solid-state
battery
charger
A constant
voltage,
Battery
charger.(fig.
A2.6-lo),
located
in the CM lower equipment bay, is incorporated
(MDC-3) controls
power
switch
into the EPS. The BATTERY CHARGERselector
as well as connecting
the charger output to the
input to the charger,
selected
battery
(fig.
A2.6-14).
When the BATTERY CHARGERselector
(Kl) is actiswitch is positioned
to entry battery
A, B, or C, a relay
vated completing
circuits
from ac and dc power sources to the battery
charger.
Battery
charger output
is also connected
to the selectedbattery
to be charged through
contacts
of the MAIN BUS TIE motor switch.
Positioning
the MAIN BUS TIE switch
(A/C or B/C) to OFF for battery
A or B,
and both switches
to OFF for battery
C will
disconnect
main bus loads from
the respective
batteries
and also complete the circuit
from the charger to
the battery.
The battery
charger is provided
25 to 30 volts
from both main!dc
All three
buses and 115 volts
400-Hz T-phase from either
of the ac buses.
de input and produce
phases of ac are used to boost the 25- to TO-volt
In addition,
phase A of the ac is used to
40 volts
dc for charging.
The logic
network in the charger,
supply power for the charger circuitry.
which consists
of a two-stage
differential
amplifier
(comparator),
Schmitt
trigger,
current
sensing resistor,
and a voltage
amplifier,
sets up the
The first
stage of the comparator
is
initial
condition
for operation.
thus setting
the Schmitt
in the on mode, with the second stage off,
Maximum base drive
trigger
first
stage to on with the second stage off.
is provided
to the current
amplifier
which turns the switching
transistor
to the on mode. With the switching
transistor
on, current
flows from the
transformer rectifier
through the switching
transistor,
current
sensing
resistor,
and switch choke to the battery
being charged. Current lags
As current
flow increases,
the
voltage
due to switching
choke action.
voltage
drop across the sensing resistor
increases,
and at a specific
level
sets the first
stage of the comparator
to off and the second stage
A-56
I_ . _--II _.-._I..
_..-4-_"1"e._.-II
_l-.l..-_-~~,-"--
-
r --w--m
i
t-[
I I
r--
1
I-
---
1
i
I
i
I[
L-
CURRENT
AMPLIFIER
I--
i
I
.----
J
DC INDICATORS
SW (MDC-3)
VOLTMETER
+
1
TELEMETRY
DC INDICATORS
SW (MDC-3)
BATTERY
CHARGE
(MDC-3)
AMMETER
_ _
C
B
.I1
:
rI
BATTERY CHARGER
BAT A CHG
(MDC-5)
*
BATTERY
BUS A
I
BAT A PWR
ENTRY/POST
(RHEB 250)
SWITCHING
DIODE
m- J
LANDING
BOA
CB41
(ON BATTERY)
40
VDC
BATTERY
MNDC
DC NEG
_-
.,
Figure
b
~2.6-ILO.- Battery
charger
A-57
block
diagram.
A
to
on.
The voltage
amplifier
is set off to reverse
the Schmitt
trigger
to first
stage off and second stage on.
This sets the current
amplifier
off,
which in turn sets the switching
transistor
off.
The switching
transistor
in the off mode terminates
power from the source,
causing the
field
in the choke to continue
collapsing,
into
the battery,
discharging
then through
the switching
diode and the current
sensing resistor
to the
opposite
side of the choke.
As the EMF in the choke decreases,
current
through
the sensing resistor
decreases,
reducing
the voltage
drop across
the resistor.
At some point,
the decrease in voltage
drop across the
sensing resistor
reverses
the comparator
circuit,
setting
up the initial
condition
and completing
one cycle of operation;.
The output load current,
due to the choke action,
remains&relatively
constant
except for the small
variation
through
the sensing resistor.
This variation
is required
to set
and reset the switching
transistor
and Schmitt trigger
through
the action
of the comparator.
Battery
charger output is regulated
by the sensing resistor
until
battery
voltage
reaches approximately
37 volts.
At this point,
the biased voltage
sensor circuit
is unbiased,
and in conjunction
with the
sensing resistor,
provides
a signal
for cycling
the battery
charger.
As
battery
voltage
increases,
the internal
impedance of the battery
increases,
decreasing
current
flow from the charger.
the battery
is
At 39.8 volts,
fully
charged and current
flow becomes negligible.
Recharging
the batbattery
amp hour input equates amp hours previously
disteries
until
charged from the battery
assures sufficient
battery
capacity
for mission
completion.
The MSFN will
monitor
this function.
If there is no contact
with the MSFN, battery
charging
is terminated
when the voltmeter
indicates
39.5 V dc with the DC INDICATORS switch set to the BAT CHARGERposition.
output
placing
Battery
Charger voltage
is monitored
is monitored
on the inner
the DC INDICATORS switch
charger current
output is
on the DC VOLTS METER (MDC-3).
Current
scale of the DC AMPS meter (MDC-3) by
(MIX-3) to the BAT CHARGERposition.
telemetered
to the MSFN.
battery
A or B, the respective
When charging
BAT RLY BUS-BAT A or B
circuit
breaker
(MDC-5) is opened to expedite
recharge.
During this
period,
only one battery
will
be powering the battery
relay bus.
Relay
bus voltage
can be monitored
by selecting
positions
4 and B on the Systems
Test Meter (LEB-101) and from the couches by the Fuel Cell-Main
Bus B-l
and Fuel Cell - Main Bus A-3 talk-back
indicators
(MDC-3) which will be
barber-poled.
If power is lost to the relay bus, these indicators
will
revert
to the gray condition,
indicating
loss of power to the relay bus
and requiring
remedial
action.
Recharge of a battery
immediately
after
it is exposed to any appreciable
loads requires
less time than recharge
of a battery
commencing
30 minutes or more after
it is disconnected
from these loads.
Therefore,
it is advantageous
to connect batteries
to the charger as soon as possible
A-58
__-.
after
they are disconnected
overall
recharge
time.
from the main buses
since
this
decreases
The dc and ac power distribution
to components
Power distribution.A singleof the EPS is provided
by two redundant
buses in each system.
point ground on the spacecraft
structure
is used to eliminate
ground loop
Sensing and control
circuits
are provided
for monitoring
and
effects.
protection
of each system.
Distribution
of dc power (fig.
~2.6-11) is accorr,plished with a twowire system and a series
of interconnected
buses, switches,
circuit
The dc negative
buses are connected to
breakers,
and isolation
diodes.
The buses consist
of the following:
the vehicle
ground point
(VGP).
a.
and/or
Two main de buses
entry and postlanding
(A and B),
batteries
powered by the
A, B, and C.
three
fuel
cells
Two battery
buses (A and B), each powered by its respective
b.
C can power either
entry and postlanding
battery
A and B. Battery
both buses if batieries
A and/or B fail.
Flight
and postlanding
bus,
or directly
by the three
and diodes,
B, and C, through dual diodes.
C.
d.
diodes.
e.
f.
through
Flight
bus,
Nonessential
Battery
relay
the individual
powered
bus,
through
powered
p owered through both
entry and postlanding
both
through
main dc buses
either
or
main dc buses
batteries
A,
and isolation
dc main bus A or B.
bus, powered by entry and postlanding
battery
buses and isolation
diodes.
batteries
Pyro b?Lzes, isolated
from the main electrical
power
system
when
gA capbility
is provided
to connect either
powered by the pyre batteries,
entry battery
to the A or B pyro system in case of loss of a pyro battery.
completely
isolated
from the main
SM jei;tison
controllers,
h.
electrical
power system until
activated
during CSM separation,
after
which they are powered by the fuel cells.
Power from the fuel cell power plants
can be connected to the main
de buses through
six
motor
switches
(part of overload/reverse
current
circuits
in the SM) which are controlled
by switches
in the CMlocated
to either
or both of the main
on MDC-3. Fuel cell power can be selected
de buses.
Six talk-back
indicators
show gray when fuel cell output
is
When an overload
condition
connected and striped
when disconnected.
the overload-reverse
current
circuits
in the SM automatically
occurs,
A-59
I I-.
_.__I..
_“.,_,..-_.___._._...._
____I
__-__---.-_.I-__..--
I^_I-^.
.-_---.-___.-.-
disconnect
the fuel cell power plants
from the overloaded
bus and provide
visual
displays
(talk-back
indicator
and caution
and warning lamp illumination)(FC
BUS DISCONNECT) for isolation
of the trouble.
A reverse
current
condition
will
disconnect
the malfunctioning
power plant from the dc sysThe de undervoltage
sensing circuits
(fig.
~2.6-12) are provided to
tem.
indicate
bus low-voltage
conditions.
If voltage
drops below 26.25 volts
dc, the applicable
de undervoltage
light
on the caution
and warning panel
Since each bus is capable of handling
all EPS
(MDC-2) will
illuminate.
an undervoltage
condition
should not occur except in an isolated
loads,
if too many electrical
units
are placed on the bus simulinstance;
taneously
or if a malfunction
exists
in the EPS. A voltmeter
(MDC-3) is
provided
to monitor
voltage
of each main dc bus, the battery
charger,
and
each of the five batteries.
An ammeter is provided
(MDC-3) to monitor
A, B, C, and the battery
current
output of fuel cells
1, 2, 3, batteries
charger.
During high power demand or emergencies,
supplemental
power to the
main de buses can be supplied
from batteries
A and B via the battery
buses
During entry,
spacecraft
and directly
from battery
C (fig.
~2.643).
power is provided
by the three entry and postlanding
batteries
which are
connected to the main dc buses prior
to CSM separation;
placing
the
MAIN BUS TIE switches
(MDC-5) to BAT A/C and BAT B/C provides
this
function
after
closing
the MAIN A-BAT C and MAIN B-BAT C circuit
breakers
The switches
are manually
placed to OFF after
completion
of
(MB-275).
RCS purge and closing
the FLIGHT AND POST LDG-BAT BUS A, BAT BUS B, and
BAT C circuit
breakers
(RHEB-275) during main chute descent.
The AUTO
position
provides
an automatic
connection
of the entry batteries
to the
main dc buses at CSM separation.
The auto function
is used only on the
launch pad after
the spacecraft
is configured
for a LES pad abort.
~2.6-11, permits isolating
nonA nonessential
bus, as shown on fig.
essential
equipment during a shortage
of power (two fuel cell power
The flight
bus distributes
power to inflight
telecommuniplants
out).
cations
equipment.
The flight
and postlanding
bus distributes
power to
some of the inflight
telecommunications
equipment,
float
bag No. 3 conthe ECS postlanding
vent and blower control,
and postlanding
comtrols,
munications
and lighting
equipment.
In flight,
the postlanding
bus receives power from the fuel cells
and/or entry and postlanding
batteries
through
the main de buses.
After
completion
of RCS purge during main
the entry batteries
supply power to the postlanding
bus
chute descent,
directly
through
individual
circuit
breakers.
These circuit
breakers
(FLIGHT & POST LANDING-BAT BUS A, BAT BUS B, and BAT C - RHEB-275) are
normally
open in flight
and closed during main chute descent just prior
to positioning
the MAIN BUS TIE switches
to OFF.
Motor switch contacts
which close when the MAIN BUS TIE switches
are
placed to ON, complete the circuit
between the entry and postlanding
batteries
and the main dc buses, and open the connection
from the battery
A-60
The battery
relay bus provides
dc power to the
charger
to the batteries.
the fuel cell and inverter
control
circuits,
fuel cell
ac sensing units,
and the fuel cell+nain
BUS A and B talk-back
reactant
and radiator
valves,
batteries
supply power to ordnance
indicators
on MDC-3. The pyrotechnic
devices
for separation
of the LES, S-IVB, forward heat shield,
SM from
CM,
and for deployment_and
release
of the drogue and main parachutes
during
a pad abort,
high-altitude
abort,
or normal mission
progression.
The three fuel cell power plants
supply power to the SM jettison
controllers
for the SM separation
maneuver.
A2.6-14) is accomplished
Distribution
of ac power (fig.
with a fourwire system via two redundant
buses, ac bus 1 and ac bus 2. The ac neutral
bus is connected
to the vehicle
ground point.
The ac power is provided by one or two of the solid-state
ll5/200-volt
400-Hz 3-phase inverters.
The dc power is routed to the inverters
through
the main dc
Inverter
No. 1 is powered through dc main bus A, inverter
No. 2
buses.
through
dc main bus B, and inverter
No. 3 through
either
de main bus A
or B by switch selection.
Each of these circuits
has a separate
circuit
breaker
and a power control
motor switch.
Switches for applying
power
to the motor switches
are located
on MDC-3. All three inverters
are
identical
and are provided
with overtemperature
circuitry.
A light
inin the caution/warning
group on MDC-2, illuminates
at 190" to
dicator,
indicate
an overtemperature
situation.
Inverter
output
is routed through
a series
of control
motor switches
to the ac buses.
Six switches
(MDC-3)
control
motor switches
which operate contacts
to connect or disconnect
the inverters
from the ac buses.
Inverter
priority
is 1 over 2, 2 over
and
3
over
1
on
any
one
ac
bus.
This
indicates
that
inverter
2 cannot
3,
be connected
to the bus until
the inverter
1 switch
is positioned
to OFF.
Also,
when inverter
3 switch is positioned
to ON, it will
disconnect
inverter
1 from the bus before
the inverter
3 connection
will
be performed.
The motor switch circuits
are designed to prevent
connecting
two inverters to the same ac bus at the same time.
The ac loads receive
power from
either
ac bus through bus selector
switches.
In some instances,
a single
phase is used for operation
of equipment and in others all three.
Over~2.6-12) are provided
undervoltage
and overload
sensing
circuits
(fig.
for each bus.
An automatic
inverter
disconnect
is effected
during an
overvoltage.
The ac bus voltage
fail
and overload
lights
in the caution/
warning group (MDC-2) provide
a visual
indication
of voltage
or overload
malfunctions.
Monitoring
voltage
of each phase on each bus is accomplished
by selection
with the AC INDICATORS switch
(MDC-3).
Readings are
displayed
on the AC VOLTS meter (MDC-3).
Phase A voltage
of each bus is
telemetered
to MSFN stations.
Several precautions
should be taken during any inverter
switching.
The first
precaution
is to completely
disconnect
the inverter
being taken
out of the circuit
whether due to inverter
transfer
or malfunction.
The
second precaution
is to insure
that no more than one switch on AC BUS 1
or AC BUS 2 (MIX-3) is in the up position
at the same time.
These
~-64
k
,
P
1
+
INVERTER NO.
1
+
INVERTER NO.
2
+
INMRTER
3
NO.
ONLY ONE INMRI~K
CAN POWEF BUS AT
ANY ONE TIME
L
INVERTER NO.
1
4-
INVERTER NO.
2
-
INVERTER NO.
3
ONLY ONE INMRTER
CAN POWER BUS AT
ANY ONE TIME
EPS SENSOR
Ib
I
-1
AC SENSE UNIT
AND AC
INDICATORS SW
(VOLTMETER)
AC SENSE UNIT
AND
;NDI~TORS
(VOLTMETER)
sw
PUMP MOTORS - FUEL CELL 1
FUEL CELL
pH SENSOR (+A)
c
pH SENSOR (+A)
c
ptl SENSOR (CA)
pH SENSOR (‘+A)
PUMP MOTORS - FUEL CELL 2
PUMP MOTORS - F&L CELL 2
-c
pH SENSOR (d)A)
-c
pH SENSOR ( +A)
c
4
CRYOGENIC
FUEL 0TY AMPL 1 (6 C)
+
CRYOGENIC
FUEL OTY AMPL 2 (CC)
4
CRYOGENIC
FAN MOTORS - S% 1
----)
CRYOGENIC
FAN MOTORS - SYS 2
PUMP MOTORS - FUEL CELL 3
PUMP MOTORS - FUEL CELL 3
--)
BATTERY CHARGER
----,
PATTERY CHARGER
-
TELECOMMUNICATIONS
----,
TELECOMMUNICATIONS
-----,
ENVlRONLNNTAL
---,
STABILIZATION
AND
CONTROL
----,
SPS GAUGING
(CC)
SYSTEM (+
_3
EXTERIOR LIGHTING
(+A)
d
INTERIOR LIGHTING
@A, 98)
_3
GLN AC POWER ( $0)
Figure
6. *
CONTROL SYSTEM
A2.6-lb.-
Alternating
--)
ENVIRONMENTAL
CONTROL
*
STABILIZATION
AND CONTROL
+
SPS GAUGING
(+O
--k
EXTERIOR LIGHTING
(‘#B)
(#A)
4
INTERIOR LIGHTING
d
GCN AC POWER ($4)
+
ORDEAL 1661
current
SYSTEM @AA.$‘C, %?
SYSTEM
power distribution.
NOTE:
1. For complerc
AC distribution
bm&out rchr
to individwl
system section
precautions
are necessary
to assure positive
power transfer
since power
to any one inverter
control
motor switch is routed in series
through
the
A third
precaution
must be exercised
to preswitch of another
inverter.
clude a motor switch lockout
when dc power to inverter
3 is being transThe AC
ferred
from dc main bus A to dc main bus B, or vice versa.
INSERTER 3 switch
(MIX-3) should be held in the OFF position
for 1 second
operation
from one main dc bus to the
when performing
a power transfer
other.
Performance
Alternating;
current
formance and design data
Alternating
and Design
Data
and direct
current
data.for the EPS is as follows:
The ac and dc per-
current
Phases
3
Displacement
120 * 2"
Steady-state
Transient
voltage
voltage
-1.5)
115.5 (+L
3 phases)
115 (+35,
-65)
V ac (average
V ac
Recovery
To 115 * 10 V within
state within
50 ms
Unbalance
2 V ac (worst
Frequency limites
Normal (synchronized
to central
timing
equipment)
Emergency (loss
central
timing
equipment)
of
phase
400 f 3 Hz
400 f 7 Hz
Wave characteristics
(sine wave)
Maximum distortion
Highest harmonic
Crest factor
5 percent
4 percent
1.414 f 10 percent
Rating
1250 V ac
15 ms, steady
from average)
Direct
current
Steady-state
limits
Normal
voltage
29 * 2.0 V dc
26.2
V dc
Minimum CM bus
Min Precautionary
bus
CM
26.5 V dc (allows
for
cyclic
loads)
Maximum CM bus
Max Precautionary
bus
CM
31.0 v dc
30.0 V dc (allows
for
cyclic
loads)
During postlanding
preflight
checkout
periods
Ripple
and
27 to 30 V dc
1 V peak to peak
voltage
Operational
Limitations
and Restrictions
Fuel cell power plants.Fuel cell power plants
are designed to
function
under atmospheric
and high-vacuum
conditions.
Each must be able
to maintain
itself
at sustaining
temperatures
and minimum electrical
loads
at both environment
extremes.
To function
properly,
fuel cells must operate under the following
limitations
and restrictions:
External
nonoperating
temperature
-20"
Operating
temperature
inside
SM
+30" to 145" F.
External
Atmospheric
nonoperating
to +140" F.
pressure
Normal
27 to 31 V dc
voltage
Minimum operating
voltage
at terminals
Emergency operation
20.5 V dc at 2295 watts
(gross
Power
level)
Normal
.
._
.._-.,
.._
.
..-
_~...---
.-..
-.“--_~-
operation
v ac
27
-“pi._._
..---.-
^
_*.l
“~-
____il_-
Maximum operating
at terminals
Fuel cell
overload
voltage
31.5 V dc
disconnect
Maximum reverse
75 amperes no trip,
112 amperes
disconnect
after
25 to 300 seconds
current
1 second minimum before
disconnect
Minimum sustaining
power/ 420 watts
fuel cell power plant
(with in-line
heater OFF)
In-line
heater power
(sustain
F/C skin temp
above 385" F min)
160 watts
(5 to 6 amps)
Maximum gross power
under emergency
conditions
2295 watts
Nitrogen
pressure
50.2
Reactant
Oxygen
pressure
at 20.5
to 57.5 psia
58.4 to 68.45
nominal)
Hydrogen
57.3 to 67.0
nominal)
psia
psia
V dc min.
(53 psia,
(62.5
(61.5
psia,
psia,
Reactant
consumption/fuel
cell power plant
Hydrogen
Oxygen
PPH = Amps x (2.57
PPH = Amps x (2.04
Minimum skin temperature
for self-sustaining
operation
+385” F
Minimum skin
for recovery
temperature
in flight
+360° F
Maximum skin
temperature
+500" F
Approximate
external
environment
temperature
range outside
SC (for
radiation)
-260”
A-68
to +400" F
nominal)
x lo-')
x 10m2)
Fuel cell power plant
normal operating
temperature
range
+385” to +450" F
Condenser
operating
+150"
Purging
exhaust normal
temperature
nominal
frequency
to +175" F
Dependent on mission
and reactant
purity
O2 purge
duration
2 minutes
H2 purge
duration
80 seconds
Additional
flow
while purging
Wwn
Hydrogen
load
after
profile
tank fill
rate
Up to 0.6 lb/hr
Up to 0.75 lb/hr
( nominal
0.67
lb/hr)
The cryogenic
storage
subsystem must
Cryogenic
storage subsystem.be able to meet the following
requirements
for proper operation
of the
fuel cell power plants
and the ECS:
Minimum usable
Oxygen
Hydrogen
Temperature
fill
Oxygen
Hydrogen
quantity
at time
probe
each tank
(min)
(min)
of
-297"
-423"
Operating
pressure
@men
Normal
Minimum
Hydrogen
Normal
Minimum
Temperature
Oxygen
Hydrogen
320 lbs
28 lbs each tank
F.
F.
(approx.)
(approx.)
range
865 to 935 psia
150 psia
225 to 260 psia
100 psia
range
-325”
to
-425"
to -200"
+80° F
F
Maximum allowable
difference
in quantity
balance between tanks
A-69
,.
. - - _..-
,. ._..-.-.-.
..___I"._.-__"_ _I___,___cI_
--~..-~~-----~~~~-
p__,p_--
._._._I__I- .I...__--.
Oxygen tanks
Hydrogen
and 2
No. 1 and
tanks
2 to
3 percent
No. 1
Pressure
relief
valve
operation
Crack pressure
Oxygen
Hydrogen
Reseat pressure
Oxygen
Hydrogen
Full flow, maximum
relief
Q-vge n
Hydrogen
4 percent
983 psig
273 psig
min.
min.
965 psig min.
268 psig min.
1010 psig max.
285 psig max.
Additional
data about 1imitatiOnS
and restrictions
Additional
data.may be found in the CSM/LM Spacecraft
Operational
Data Book SNA-8-D-027,
vol I, (CSM ~1168-447).
Systems Test
Meter
The SYSTEMS TEST meter and the alphabetical
and numerical
switches
located
on panel 101 in the CM LEB, provide
a means of monitoring
vario&s
measurements within
the SC, and verifying
certain
parameters
displayed
only by event indicators.
The following
can be measured using the
SYSTEMS TEST meter, the respective
switch positions,
and the range of
each sensor.
Normal operating
parameters
of measurable
items are covered
in the telemetry
listing.
Conversion
of the previously
listed
measurements to the SYSTEMS TEST
meter indications
are listed
in Table A2.6-IV.
The XPNDR measurements
are direct
readouts
and do not require
conversion.
A-70
.l_l - -
~~-I- . . ~"--..
"l.--__l^.-..
_-.._-L;-
__l_-.,_l
TABLE &Z.&III.-
Systems test
indicatioafid(";yT
SYSTEMSTEST DATA
>sitions
Alphabetical
select
Switch
Numerical
select
identity
Sensor range
0 to 75 psia
v2 pressure, psia
F/C 1 SC 206OP
F/C 2 SC 2061~
F)C 3 SC 2062~
A
B
C
1
1
1
0
o2 pressure, psia
F/C 1 SC 2066~
F/C 2 SC 2067P
F'/C 3 SC 2068P
1
2
2
D
A
B
92 pressure, psia
F/C 1 SC 2069~
F/C 2 SC 2070P
F'/C 3 SC 2071P
2
2
3
C
D
A
3
3
3
B
C
D
<pS radiator
outlet
F/C 1 SC 208~
F'/C 2 SC 20881
F~C 3 SC 2089~
75 psia
t0
0 to 75 Pia
-50" to t300" F
temperature
3attery manifold
pressure, psia
A
0
Bat7; relay
B
o to +45 V dc
D
0 to t10 amps
5
A
o to +200e F
6
5
B
D
C
D
A
C
bus CCO232V
LM power
SpS oxidizer
SP 0049T
line
temperature
t0
20
psia
-50" to ~50" F
CM-RCS oxidize~y~a~eCRte2;l~Tature
-P engine,
+Y engine, sys B CR 2116~
-P engine, sys B CR 2110T
CWengine, sys B CR 2119T
CCWengine, sys A CR 2114T
-Y engine, sys A CR 2103T
2
6
6
pwr output
XPNDR
A
>I.0 V dc (nornina
AGC signal
XPNDR
B
Test A.0 V ac
Operate 0.0 to
4.5 v ac
Phase lockup
XPNDR
C
Locked A.0 V
Unlocked 4.8
NOTE:
position
7 on the numerical
selector
switch
A-
_.---.c_---
__^__1__--
.
is an off Fosltlon*
71
___---..----------
-___
.---
Command Module
Interior
The command module interior
lighting
illumination
for activities
in the couch,
panel lighting
to
areas, and back-lighted
and switch positions.
Tunnel lighting
is
concerned with LM activity.
Lighting
system (fig.
A2.6-15)
furnishes
lower equipment bay and tunnel
read nomenclature,
indicators,
provided
on SC which will
be
Floodlighting
for illumination
of work areas is provided
by use of
fluorescent
lamps.
Integral
panel and numerics lighting
is provided
by
electroluminescent
materials.
Tunnel lights
are incandescent.
Pen
flashlights
are provided
for illuminating
work areas which cannot be
illuminated
by the normal spacecraft
systems, such as under the couches.
Electroluminescence
(EL) is
from a crystalline
phosphor
(Z#)
the phenomena whereby light
is emitted
placed as a thin layer between two
closely
spaced electrodes
of an electrical
capacitor.
One of the electrodes is a transparent
material.
The light
output varies
with voltage
and frequency
and occurs as light
pulses,
which are in-phase
with the
input frequency.
Advantageous
characteristics
of EL for spacecraft
use
are an "after-glow"
of less than 1 second, low power consumption,
and
negligible
heat dissipation.
RAL AND NUMERICS (PANEL tlctm
FLOODLIGHT FIXTURES
Figure
~2.6-15.-
CM interior
A-72
1.ighting.
TABLE .42.6-m.EPS
radiator
Outlet
temperature
(" F.)
Systems
test
meter
display
0.0
0.2
0.4
0.6
0.8
1.0
;
6
9
12
15
1.2
1.4
1.6
1.8
2.0
18
21
24
27
30
2.2
2.4
2.6
2.8
3.0
529
45
,':c
3.6
48
51
54
43::
2
4.2
63
t::
4.8
5.0
6”:
72
75
33
36
-50
-36
-22
-8
SYSTEMSTEST METER INDICATIONS
CM-R%
Oxidizer
valve
tawerature
(” F.)
LM
Power
(amps)
-50
-46
-4;
-33
+6
+20
-34
-30
+34
+4a
+62
+76
+90
-26
-22
-18
-14
-10
0
+14
+18
+22
+26
+30
6.4
6.8
7.2
7.6
8.0
4.4
4.8
5.2
8.4
8.8
9.2
9.6
10.0
Battery
manifold
Pressure
@=A)
Batter:
relay
bus
('J dc
0.00
0.80
1.60
2.40
3.20
4.00
0
1.8
4.80
5.60
6.40
7.20
8.00
5::
+174
+188
+202
+216
+230
+46
t50
4’:
2.4
2.8
3.2
21:
+4-i?
16
24
0.8
1.2
1.6
2.0
1;
0
+4
+10
t34
+38
8”
0.4
+104
+118
+I32
t146
+16o
+244
1258
+272
e86
+300
SPS
temperature
(" F.)
88
96
104
112
120
128
136
144
152
160
x8
176
184
192
200
8.80
9.60
10.40
11.20
12.00
12.80
13.60
14.40
15.20
16.00
16.80
17.60
18.40
19.20
20.00
3.6
5.4
7.2
9.0
lo.8
12.6
14.4
16.2
18.0
19.8
21.6
23.4
25.2
27.0
28.8
30.6
32.4
34.2
36.0
37.8
39.6
41.4
43.2
45.0
The interior
floodlight
system consists
of six
Floodlight
system.floodlight
fixture
assemblies
and three control
panels (fig.
~2.6-16).
Each fixture
assembly contains
two fluorescent
lamps (one primary
and
one secondary)
and converters.
The lamps are powered by 28 V dc from main
This assures a power source for lights
de buses A and B (fig.
A2.6-17).
in all areas in the event either
bus fails.
The converter
in each flood28 V dc to a high-voltage
pulsating
dc for operation
light
fixture
converts
of the fluorescent
lamps.
the left
Floodlights
are used to illuminate
three specific
areas:
the right
main display
console,
and the lower
main display
console,
control
of lighting
of the
Switches on MDC-8provide
equipment bay.
left
main display
console area.
Switches on MDC-5provide
control
of
Switches for control
lighting
of the right
main display
console area.
of lighting
of the lower equipment bay area are located
on LEB-100.
Protection
for the floodlight
circuits
is provided
by the LIGHTING - MN A
and MN B circuit
breakers
on FXEB-226.
Each control
panel has a dimming (DIM-l-2)
toggle switch control,
a
and an on/off
(FIXED-OFF) toggle
switch
rheostat
(FLOOD-• FF-BRT) control,
control.
The DIM-l position
provides
variable
intensity
control
of the
primary
flood lamps through
the FLOOD-OFF-BRT rheostat,
and on-off
control
The DIM-2 position
of the secondary lamps through the FIXED-OFF switch.
provides
variable
intensity
control
of the secondary lamps through
the
and on-off
control
of the primary
lamps through
FLOOD-OFF-BRT rheostat,
When operating
the primary
lamps under variable
the FIXED-OFF switch.
turn on of the lamps is acquired
after
intensity
control
(DIM-l position),
In transferring
the FLOOD-OFF-BRT rheostat
is moved past the midpoint.
variable
intensity
control
to the secondary lamps, the FLOOD-OFF-BRT
rheostat
should first
be rotated
to the OFF position
before placing
the
The rheostat
is then moved to the full
DIM switch to the DIM-2 position.
bright
setting
and should remain in this position
unless dimming is desired.
Dimming of the secondary flood lamps should not be used unless
Dimming of
dimming control
of the primary
floodlights
is not available.
the secondary lamps results
in approximately
a go-percent
reduction
in
lamp life.
The range of intensity
variation
is greater
for the primary
than the secondary floodlights.
The commander's control
panel (MIX-~) has a POST LANDING-OFF-FIXED
switch which connects the flight
and postlanding
bus to his floodlights
The POST LANDING position
provides
single
intensity
(fig.
A2.6-17).
lighting
to the commander's primary
or secondary lamps as selected
by the
It is for use during
the latter
DIM-l or DIM-2 position,
respectively.
stages of descent after main dc bus power is disconnected,
and during
postlanding.
A-74
_ .__._...-...a--
=.-
-FLOOOT
DIM FIXED
1
COUCH LIGHT
ASSEMBLIES
2
MDC-5
POST LDG
-
__^.._
COMPONENTS
6 LIGHT ASSEMBLIES
3 CONTROLPANELS
LH SIDE DISPLAY MDC-8
RH SIDE DISPLAY MDC-5
LEB 100
3 CIRCUIT BREAKERS
RHEB 226
Figure
A2.6-16.-
CM floodlight
configuration.
+ 7
I
POST LANDING BUS
+
DC MAIN BUS B
f
DC MAIN BUS A
RHEB 226
I-- _-------------------------------,
1 I
9)
I
FLOOD, MNA $
---------------------T
-
I
i
7.5A
i
-___i
t
1
DIM
II
MDC 8
t I
I
+
I
I
I
II
I
,
P
LIGHT
I
0
NEGATIVE DC BUS
,
:----~~~~-~~~E!!---- c------ I I---C---- ---!E-----J
I
63
Figure
I
I
LIGHT
~2.6~17.- CM floodlight
1
LIGHT
I
NEGATIVE DC BUS
______
I
I
L--=+-t
q
system
P
schematic.
NEGATIVE DC BUS
Integral
lighting
system.The integral
lighting
system controls
the
EL lamps behind the nomenclature
and instrument
dial faces on all MDC!
and on specif'ic
panels in the lower equipment bay, left
hand
panels,
~2.6-18
and
~2.6-19).
and
right
hand
equipment
bay
(figs.
equipment bay,
The controls
(fig.
~2.6-18) are rotary switches controlling
variable
transformers
powered through the appropriate
ac bus.
Each rotary
control
switch has a mechanical
stop which prevents
the switch being positioned
of a circuit
because of malfunctions
is performed by
to OFF. Disabling
opening the appropriate
circuit
breaker
on RHEB-226.
The INTEGRAL switch
the lighting
of panels viewed by the commander, MD&l,
on MIX-~ controls
and
the
left
half
of 2. The INTEGRAL switch on MDC-5 con7, 8, 9, 15,
trols
the lighting
of panels viewed by the LM pilot,
MDC-3, 4, 5 and 6,
16, RHEB-229 and 275, and the right half of MIX-2.
The INTEGRAL switch
on LEB-100 controls
the lighting
of MDC-10, LEB-100, 101, 122 and the
of the lighting
DSKY lights
on 140, RHEB-225, 226 and LHEB 306. Intensity
can be individually
controlled
in each of the three areas.
Numerics lighting
system.Numerics lighting
control
is provided
over all electroluminescent
digital
readouts.
The NUMERICS rotary
switch
the off/intensity
of numerals on the DSKY and Mission
on MDC-8 controls
Timer on MIX-2, and the range and delta V indicators
of the Entry Monitor
the off/intensity
of the
System of MDC-1. The switch on LEB-100 controls
numerals on the LEB-140 DSKY and the Mission
Timer on LHEB-306.
Protection for the integral
and numerics circuits
is provided
by the LIGHTINGNUMERICS/INTEGRAL-LEB AC 2, L MIX AC 1, and R MIX AC 1 circuit
breakers
on RHEB-226.
These circuit
breakers
are used to disable
a circuit
in
case of a malfunction.
The L MDC AC 1 circuit
breakers
also feed the
EMS roll
attitude
and scroll
incandescent
lamps.
The six light
fixtures
in the CM tunnel
provide
Tunnel lighting.illumination
for tunnel
activity
during docking and undocking.
Each of
28 V dc
the fixtures,
containing
two incandescent
lamps, is provided
through a TUNNEL LIGHTS-OFF switch on MIX-2 (fig.
~2.6-20).
Main dc bus A
distributes
power to one lamp in each fixture,.and
main dc bus B to the
other lamp.
Protection
is provided
by the LIGHTING/COAS/TUNNEL/RNDZ/
SPOT MN A and MN B circuit
breakers
on RKEB-226.
A-77
-.--.
.
_.._.
_-
-_
..L--.
-.
.:.:.>:
cls;$;:1Nj’EGRAL
.-.
B-: NUMERICS
INTERIOR LIGHTS
INTEGRAL
(MDC 81
LED LIGHTS
FLOOD
r
NUMRICS
OFF
BIT
OFF
RRT
t
INTEGRAL
OFF
BRT
(LEB 100)
Figure
~2.6-18.-
CM integral/numerics
illuminat
,ion
system.
LIGHTING NUhlERlCS'lNTCGRAL
+?BUSl
< P AC NEUT
I
hlECHANICA1
STOP
110 PLACES,
L
INTERIOR LIGHTS
lRHEB 2261
dB)
NOMFN
'
p )
I
I
QA 7)
+AC,NEUl
I
tEB-A?$
I
-
BUS 2
I
I
I
)
INTERIOR
LIGHTS
NOMEN
,
MDC 5
LHEB 306
EMS (INCANDESCENT1
hlDC 1
1 MDC 3
I
L
Figure
A2.6-lg.-
Integral
and numerics
panel
lighting
ROLL ATT IND 141
SCROLL
(6)
schematic.
I
r
---
1
I
I
I
t
I
I
6
TUNNEL
LIGHTS
--c-
I
MN B
INITIATORI
I
I
I
I
I
I
DC I
MN
BUS A I
T
I
I
I
I
r -- iI
I
I
I
MN DC1
BUS B
I
o+
TUNNEL
LIGHT
I
ASSEMBLIES
CB28
1 LH COAS 1
Figure
I
i
i
I
1
I
I
I
1
I
~2.6-20.-
II
DC NEG
Tunnel
lighting
schemat .c.
PART A2.7
ENVIRONMENTAL CONTROL SYSTEM
Introduction
The environment
control
system (ECS) is designed to provide
the
flight
crew with a conditioned
environment
that is both life-supporting,
The ECS is aided in the accomplishment
and as comfortable
as possible.
of this task through
an interface
with the electrical
power system,
The ECS also interfaces
with
which supplies
oxygen and potable
water.
the electronic
equipment
of the several
Apollo systems,
for which the
with the lunar module (LM) for pressurizing
ECS provides
thermal
control,
the LM, and with the waste management system to the extent
that the
water and the urine dump lines
can be interconnected.
The ECS is operated
continuously
During this operating
period
phases.
three major functions
for the crew:
a.
Spacecraft
atmosphere
b.
Water management
c.
Thermal
throughout
all Apollo mission
the system provides
the following
control
control.
Control
of the spacecraft
atmosphere consists
of regulating
the
pressure
and temperature
of the cabin and suit gases; maintaining
the
desired humidity
by removing excess water from the suit and cabin gases;
controlling
the level of contamination
of the gases by removing CO*,
and ventilating
the cabin after
landing.
matter;
odors, and particulate
There are provisions
for pressurizing
the lunar module during docking
and subsequent
CSM/LM operations.
Water management consists
of collecting,
sterilizing,
and storing
and delivering
chilled
and
the potable water produced in the fuel cells,
heated water to the crew for metabolic
consumption,
and disposing
of the
excess potable
water by either
transferring
it to the waste water system
Provisions
are also made for the collection
or by dumping it overboard.
and storage of waste water (extracted
in the process of controlling
humidity),
delivering
it to the glycol
evaporators
for supplemental
and dumping the excess waste water overboard.
cooling,
A-81
Thermal control
consists
of removing the excess heat generated
by
the crew and the spacecraft
equipment,
transporting
it to the cab heat
exchanger
(if required)
, and rejecting
the unwanted heat to space,
either
by radiation
from the space radiators,
or in the form of steam
by boiling
water in the glycol
evaporators.
the
Five subsystems operating
required
functions:
a.
Oxygen subsystem
b.
Pressure
c.
Water subsystem
d.
Water-glycol
e.
Postlanding
suit
circuit
in
conjunction
with
each other
provide
(PSC)
subsystem
ventilation
(PLV) subsystem.
The oxygen subsystem controls
the flow of oxygen within
the command
module (CM); stores a reserve
supply of oxygen for use during entry
and emergencies;
regulates
the pressure
of oxygen supplied
to the subsystem and PSC components;
controls
cabin pressure
in normal and
emergency (high flow-rate)
modes; controls
pressure
in the water tanks
and glycol
reservoir;
and provides
for PSC purge via the DIRECT 02
valve.
The pressure
suit circuit
provides
the crew with a continuously
conditioned
atmosphere.
It automatically
controls
suit gas circulation,
pressure,
and removes debris,
excess moisture,
odors,
and temperature;
and carbon dioxide
from both the suit and cabin gases.
The water subsystem (potable
section)
collects
and stores potable
water;
delivers
hot and cold water to the crew for metabolic
purposes;
and augments the waste water supply for evaporative
cooling.
The waste
water section
collects
and stores water extracted
from the suit heat
exchanger,
and distributes
it to the water inflow
control
valves of the
evaporators,
for evaporative
cooling.
The water-glycol
potable
water chiller,
cooling
for the cabin
The postlanding
circulating
ambient
subsystem provides
and the spacecraft
atmosphere.
ventilation
air through
cooling
for
equipment;
subsystem provides
the command module
a-82
the PSC, the
and heating
or
a means for
cabin after
landing.
-
Functional
Description
The environmental
control
system operates
continuously
throughout
all mission phases.
Control
begins during preparation
for launch and
continues
through
recovery.
The following
paragraphs
describe
the
operating
modes and the operational
characteristic
of the ECS from the
time of crew insertion
to recovery.
Spacecraft
atmosphere control.SUIT CIRCUIT RETURN VALVE is closed;
During prelaunch
operations
and the DIRECT O2 valve is
the
opened
slightly
(approximately
0.2 pound per hour flowrate)
to provide
an
oxygen purge of the PSC. Just before prime crew insertion
the 02
flowrate
is increased
to 0.6 pound per hour.
This flow is in excess of
that required
for metabolic
consumption
and suit leakage.
This excess
flow causes the PSC to be pressurized
slightly
above the CM cabin.
The
slight
overpressure
maintains
the purity
of the PSC gas system by
preventing
the cabin gases from entering
the PSC.
Any changes made in the pressure
or composition
of the cabin gas
during the prelaunch
period
is controlled
by the ground support
equipment
through the purge port in the CM side hatch.
As soon as the crew connects into the PSC, the suit gas becomes
contaminated
by C02, odors, moisture,
and is heated.
The gases are
circulated
by the suit compressor through
the CO2 and odor absorber
assembly where a portion
of the CO2 and odors are removed; then through
the heat exchanger,
where they are cooled and the excess moisture
is
Any debris that might get into the PSC is trapped by the
removed.
debris trap or on felt pads on the upstream side of each LiOH cartridge.
During the ascent,
the cabin remains at sea level pressure
until
the ambient pressure
decreases a nominal 6 psi.
At that point the CABIN
PRESSURE RELIEF valve vents the excess gas overboard,
maintaining
cabin
pressure
at 6 psi above ambient.
As the cabin pressure
decreases,
a
relief
valve in the O2 DEMAND REGULATORvents suit gases into the cabin
to maintain
the suit pressure
slightly
above cabin pressure.
Sometime after
attaining
orbit
it will
be necessary
to close the
DIRECT 02 valve to conserve oxygen.
(Refer to Volume 2, Apollo Operations
Handbook for the procedure.)
After the DIRECT O2 valve is closed,
makeup oxygen for the PSC is supplied
by the DEMANDREiGULATORwhen the
SUIT CIRCUIT RETURN VALVE is closed or from the cabin via the cabin
pressure
regulator
when the SUIT CIRCUIT RETURN VALVE is open.
A-83
Before changing
from a suited to a shirtsleeve
environment
it is
necessary
to open the SUIT CIRCUIT RETURN VALVE, for the following
reasons.
When a suit is vented (by removing helmet,
gloves,
etc.)
some
of the PSC gases flow into the cabin, which results
in contaminating
the
cabin gas, and in lowering
suit pressure
relative
to cabin pressure.
Opening the SUIT CIRCUIT RETURN VALVE allows cabin gas to circulate
through
the PSC for scrubbing,
and tends to equalize
the pressure
differential
between the PSC and cabin.
If the valve is not opened,
the resultant
pressure
differential
will
cause the suit DEMANDREG
to dmp oxygen into the PSC at a flowrate
that will
turn on the 02
Opening the SUIT CIRCUIT RETURN VALVE will
FLOW HI warning light.
correct
this situation.
During normal space operations,
the cabin pressure
is maintained
at a nominal 5 psia by the cabin pressure
regulator,
at flowrates
up to
1.4 pounds of owgen per hour.
In the event a high leak rate develops,
supply
oxygen at high flow
the EMERGENCYCABIN PRESSURE regulator
will
rates to maintain
the cabin pressure
above 3.5 psia for more than 5 minutes, providing
the leak is effectively
no larger
than a l/2-inch
hole.
When performing
depressurized
operations
the suit circuit
is maintained
above 3.5 psia by the O2 DEMANDREGULATOR; the
pressure
regulator
shuts
off
automatically
to prevent
wasting
pressure
cabin
oxygen.
Prior to entry SUIT CIRCUIT RETURN VALVE is closed,
isolating
the suit circuit
from the cabin; the O2 DEMAND REGULATORthen controls
Cabin pressure
is maintained
during the descent by
suit pressure.
the cabin pressure
regulator
until
the ambient pressure
rises to a
maximum of 0.9 psi above cabin pressure.
At that point the cabin relief
valve will
open, allowing
ambient air to flow into the cabin.
As the
cabin pressure
increases,
the 02 DEMANDREGULATOR admits oxygen into
the suit circuit
to maintain
the suit pressure
slightly
as measured at the suit compressor inlet
manifold.
below
the
cabin,
After spacecraft
landing,
the cabin is ventilated
with ambient air
by postlanding
ventilation
fan and valves.
When the CM is floating
upright
in the water,
the POST LANDING VENT switch is placed in the HIGH
(day) or LOW (night)
position.
Either
of these positions
will
supply
power to open both vent valves and start
the fan.
In the HIGH position,
the fan will
circulate
150 cubic feet per minute (cfm); LOW, 100 cfm.
An attitude
sensing device automatically
closes both valves and removes
power from the fan motor when the CM X axis rotates
more than 60 degrees
from vertical.
Once the device is triggered,
it will
remain locked up
until
the CM is upright,
and the POST LANDING VENT switch is placed in
This action
resets the control
circuit
for normal
the OFF position.
The PLVC switch on panel 376 provides
an override
system operation.
~-84
control
for opening the PLV valves and turning
on the fan in case the
attitude
sensor is locked up and cannot be reset;
or when the CM is
inverted
and egress must be made through
the tunnel
hatch.
In either
case the POST LANDING VENT switch must be in the LOW or HIGH position.
Water management.In preparing
the spacecraft
for the mission,
the potable
and waste water tanks are partially
filled
to insure
an
adequate supply for the early stages of the mission.
From the time
the
fuel cells
are placed in operation
until
CSM separation,
the fuel cells
replenish
the potable
water supply.
A portion
of the water is chilled
and made available
to the crew through
the drinking
fixture
and the
food preparation
unit.
The remainder
is heated,
and is delivered
through
a separate
valve on the food preparation
unit.
the crew connects into the suit circuit
until
entry,
From the time
the water accumulator
pumps are extracting
water from the suit heat
exchanger
and pumping it into the waste water system.
The water is
delivered
to the glycol
evaporators
through
individual
water control
Provision
is made for dumping excess waste water manually
valves.
when the tank is full.
A syringe
injection
injection
of bactericide
system
to kill
is incorporated
to provide
for
bacteria
in the potable
water
periodic
system.
Thermal control.Thermal control
is provided
by two water-glycol
coolant
loops (primary
and secondary).
During prelaunch
operations
ground servicing
equipment cools the water-glycol
and pumps it through
the primary
loop, providing
cooling
for the electrical
and electronic
equipment,
and the suit and cabin heat exchangers.
The cold water-glycol
is also circulated
through
the reservoir
to m&e available
a larger
quantity
of coolant
for use as a heat sink during the ascent.
Additional
heat sink capability
is obtained
by selecting
maximum cooling
on the
CABIN TEMP selector,
and placing
both cabin fans in operation.
This
cold soa?ss the CM interior
structure
and equipment.
Shortly
before
launch,
one of the primary pmps is placed in operation,
the pump in
the ground servicing
unit is stopped,
and the unit is isolated
from the
spacecraft
system.
During the ascent,
the radiators
will
be heated by aerodynamic
friction.
To prevent
this heat from being added to the CM thermal
load,
the PRIMARY GLYCOL TO RADIATORS valve is placed in the PULL TO BYPASS
The coolant
then
position
at approximately
75 seconds before launch.
circulates
within
the CM portion
of the loop.
The heat that is generated
in the CM, from the time that the ground
servicing
unit is isolated
until
the spacecraft
reaches 1lOK feet,
is
Above 1lOK feet
absorbed by the coolant
and the prechilled
structure.
~-85
I____
_
..--
-....
.-.
_“.^_.---_.
~._.
Il------l_l
_-.-
,.--”
___-_..--._
.~..--
__111-__.“.“-11141-.1-
it is possible
primary
glycol
to reject
evaporator.
the
excess
heat
by evaporating
water
in the
After
attaining
orbit
the reservoir
is isolated
from the loop to
maintain
a reserve
quantity
of coolant
for refilling
the primary
loop
in case of loss of fluid
by leakage.
The PRIMARY GLYCOL TO RADIATORS
valve is placed in the position
(control
pushed in) to allow circulation
through
the radiators
end the radiator
outlet
temperature
sensors.
If
the radiators
have cooled sufficiently
(radiator
outlet
temperature
is
less than the inlet)
they will
be kept on-stream;
if not, they will
be
After
the
radiators
bypassed until
sufficient
cooling
has tsken place.
have been placed on-stream,
the glycol
temperature
control
is activated
(GLYCOL EVAP TEMP IN switch in AUTO); and the CABIN TEMP selector
is
positioned
as desired.
The primary
loop provides
thermal
control
throughout
the mission
unless a degradation
of system performance
requires
the use of the
secondary
loop.
Several hours before CM-SM separation
the system valves are
positioned
so that the primary
loop provides
cooling
for the cabin heat
and the suit heat exchanger.
exchanger,
the entire
cold plate network,
The CABIN TEMP control
valve is placed in the MAX COOL position,
and
both cabin fans are turned on to cold-soak
in the CM interior
structure.
Prior to separation
t'ne PRIMARY GLYCOL TO RADIATORS, and the
GLYCOL TO RADIATORS SEC valves are placed in the BYPASS position
to
From that time
prevent
loss of coolant when the CSM umbilical
is cut.
(until
approximately
1lOK feet spacecraft
altitude)
cooling
is provided
by water evaporation.
Oxygen Subsystem
The oxygen subsystem shares the oxygen supply with the electrical
power system.
Approximately
640 pounds of oxygen is stored in two
cryogenic
tanks located
in the service
module.
Heaters within
the tanks
pressurize
the oxygen to 900 psig for distribution
to the using equipment.
Oxygen is delivered
to the command module through two separate
each of which is connected to an oxygen inlet
restrictor
supply lines,
assembly.
Each assembly contains
a filter,
a capillary
line,
and a
The filters
provide
final
filtration
of gas
spring-loaded
check valve.
which are wound around the hot glycol
entering
the CM. The capillaries
they restrict
the total
O2 flow rate to a
line serve two purposes;
maximum of 9.0 pounds
CM. The check valves
per hour, and they heat the oxygen entering
serve to isolate
the two supply lines.
A-86
the
-
Downstream of the inlet
check valves the two lines tee together
and a single
line is routed to the OXYGEN-S/M SUPPLY valve on panel
326.
This valve is used in flight
as a shutoff
valve to back up the
inlet
check valves during entry.
It is closed prior
to CM-SM separation.
PART A2.8
TELECOMMUNICATIONS SYSTEM
Introduction
The communications
subsystem is the only link between the spacecraft
and the manned space flight
network (MSFN).
In this capacity,
the
communications
subsystem provides
the MSFN flight
controllers
with data
through
the pulse code modulated
(PCM) telemetry
system for monitoring
spacecraft
parameters,
subsystem status,
crew biomedical
data, event
occurrence,
and scientific
data.
As a voice link,
the communications
subsystem gives the crew the added capability
of comparing and evaluating data with MSFN computations.
The communications
subsystem,
through
its MSFN link,
serves as a primary means for the determination
of spacecraft
position
in space and rate of change in position.
CM-I.&l rendezvous
is facilitated
by a ranging
transponder
and an active
ranging
system.
Through the use of television
camera, crew observations
and public
information can be transmitted
in realtime
to MSFN. A means by which CM and
IM telemetry
and voice can be stored in the spacecraft
for later
playback, to avoid loss because of an interrupted
communications
link,
is
provided
by the communications
subsystem in the form of the data storage
equipment
(DSE).
Direction-finding
aids are provided
for postlanding
location
and rescue by ground personnel.
The following
a.
Provide
(1)
and in
storage
(2)
orbital
list
voice
Astronauts
summarizes
the
communication
via
the
telecormn
functions:
between:
intercom
CSM and MSFN via the unified
S-band equipment
and recovery
phases via the VHF/AM
(3)
CSM and extravehicular
(4)
CSM and L&l via
VHF/AM
(5)
CSM and launch
control
(6)
CSM and recovery
(7)
Astronauts
equipment
astronaut
force
and the voice
A-88
.-I~ .._-. .-_.... -x ,...-._-
general
center
(EVA) via
(LCC) via
swimmers via
log via
(USBE)
VHF/AM
PAD COMM
swimmers umbilical
intercomm
to the
data
b.
C.
Provide
three
d.
e.
f.
systems
to the MSFN of:
(1)
CSM system
(2)
Astronaut
biomedical
(3)
Astronaut
activity
(4)
EVA personal
(5)
I.&4 system
Provide
update
(1)
Digital
(2)
Digital
(WE)
equipment
in
data
status
via
life
television
support
status
system
recorded
reception
for
time-referencing
on CSM data
the
data
for
status
equipment
of:
the
computer
central
switching
(CMC)
timing
functions
between:
MSFN and CSM via
(2)
LM and CSM via
the
(3)
CSM and LM via
the VHF/AM ranging
a recovery
storage
perform
(1)
Provide
and biomed
command module
commands to remotely
ranging
(PISS)
and processing
information
(3)
Real-time
CM systems
Facilitate
status
the
USBE transponder
rendezvous
aid VHF for
Provide
a time reference
for all
except the guidance and navigation
Functional
radar
spacecraft
transponder
(RRT)
system
location.
time-dependent
subsystem.
spacecraft
sub-
Description
The functional
description
of the telecommunications
system is
divided
into four parts:
intercommunications
equipment,
data equipment,
radio frequency
equipment,
and antenna equipment.
All of these functional
groups of equipment interface
with each other to perform the system
In the functional
descriptions
of these parts,
such interfaces
tasks.
will
be apparent.
-
PART A2.9
SEQUENTIAL SYSTEMS
Introduction
Sequential
systems include
certain
detection
and control
subsystems
of the launch vechicle
(LV) and the Apollo spacecraft
(SC).
They are
utilized
during launch preparations,
ascent,
and entry portions
of a
. .
early mission
terminations,
docking maneurmssion,
preorbital
aborts,
Requirements
of the sequential
sequences.
vers, and SC separation
Figure AZ!.%1
systems are achieved by integrating
several
subsystems.
illustrates
the sequential
events controlsubsvstem
(SECS), which is the
and its interface
with the following
nucleus
of sequential
systems,
subsystems and structures:
a.
Displays
and controls
b.
Emergency
c.
Electrical
d.
Stabilization
e.
Reaction
f.
Docking
g*
Telecommunications
h.
Earth
1.
Launch
J-
Structural
detection
power
(EDS)
(EPS)
and control
control
(RCS)
(DS)
lauding
(T/C)
(ELS)
escape
(LES)
Sequential
which
(SCS)
Events
Control
Subsystem
The SECS is an integrated
subsystem consisting
may be categorized
in seven classifications
a.
Two master
b.
Two service
events
module
sequence
controllers
jettison
controllers
of 12 controllers
listed
as follows:
(MESC)
(SMJC)
-
A-90
c.
One reaction
d.
Two lunar
module
e.
Two lunar
docking
events
f.
Two earth
landing
sequence
65. One pyro
control
system
controller
(LM) separation
continuity
(RCSC)
sequence
controllers
(LSSC)
(LDEC)
controllers
verification
controllers
(ELSC)
bGX (PCVB)
Five batteries
and three fuel cells
are the source of electrical
power.
The SMJC is powered by fuel cells;
however, battery
power is
The RCSC is powered by the fuel cells
and
used for the start
signal.
batteries.
The remaining
controllers
of the SECS are powered by batteries
exclusively.
DISPLAYS
AND
CONTROLS
EMERGENCY
DETECTION SUBSYSTEM
4
STRUCTURES
.
SEQUENTIAL
EVENTS
CONTROL
ELECTRICAL
POWER
SUBSYSTEM
SUBSYSTEM
Figure
A2.9-l.-
LAUNCH
ESCAPE
SUBSYSTEM
EARTH
LANDING
SUBSYSTEM
SECS interface.
A-91
_~ . ".
~.__
.___-".,.~
_---"
_"-
~...
....-.I_"._.__C_"-.--
~.-__.-..---_---
_-_-.-_-_ll_l
.^... -.---,
Origin
of Signals
The SECS receives
manual and/or automatic
signals
and performs
control
functions
for normal mission
events or aborts.
The manual
signals
are the result
of manipulating
switches
on the main display
console
(MDC) or rotating
the Commander's translation
hand control
counterclockwise,
which is the prime control
for a manual abort.
Automatic
abort signals
are relayed by the emergency system (EDS).
-
A-92
FART A2.10
CAUTION AND WARNING SYSTEM
Introduction
The caution
and warning system (C&WS) monitors
critical
parameters
or out-ofof most of the systems in the CM and SM. When a malfunction
tolerance
condition
occurs in any of these systems, the crew is
immediately
alerted
in order that corrective
action may be taken.
Functional
Description
Upon receipt
of malfunction
or out-of-tolerance
signals,
the C&WS
simultaneously
identifies
the abnormal condition
and alerts
the crew
Each signal will
activate
the appropriate
systems
to its existence.
The master alarm circuit
status
indicator
and a master alarm circuit.
visually
and aurally
attracts
the crew's attention
by alarm indicators
Crew acknowledgment
on the MDC and by an audio tone in the headsets.
of an abnormal condition
consists
of resetting
the master alarm circuit,
while retaining
the particular
systems status malfunction
indication.
The capability
exists
for the crew to select
several
modes of observing
and of monitoring
CM or SM
systems status
and master alarm indicators
systems.
Major
Component Subsystem
Description
The C&WS consists
of one major component, the detection
unit.
It
is located behind MDC-3, and therefore
is neither
visible
nor accessible
The balance of the system is made u-p
to the crew during the mission.
aural alerting
and associated
circuits,
and those
of visual
indicators,
Visual
switches
required
to control
the various
system functions.
indicators
include
the two uppermost fuel cell electromechanical
event
devices on MDC-3, as well as all systems status
and master alarm lights.
The detection
unit circuits
consist
of comparators,
logic,
lamp
Also incorporated
are
and a master alarm and tone generator.
drivers,
two redundant
power supplies,
a regulated
+12 and a -12 V dc for the
electronics.
Inputs to the detection
unit consist
of both analog and event-type
Alarm limits
signals.
The analog signals
are in the 0 to 5 V dc range.
for these signals
trigger
voltage
comparators,
which,
in turn,
activate
thus causing activation
of the master
logic
and lamp-driver
circuits,
A-93
^-
.__
_.
.
. .
.‘_
__.
.
__I___~
__.__-
_-
-_.-.
“--.
. .__-_..-_ms....m----
_.----
~---
.-,_,^._
-.._
I.-
alarm circuit
and tone generator,
illumination
of applicable
systems
status
lights
on MDC-2, and for certain
measurements,
activation
of
applicable
electromechanical
event indicators
on MDC-3. Several event
inputs
are monitored
by the C&WS detection
unit.
These signals
originate
from solid
state and mechanical
switch closures
in malfunction
sensing
devices.
These signals
will
directly
illuminate
applicable
system status
lights
and, through
logic
circuitry,
activate
the master alarm lights
and tone generator.
One event signal,
originating
within
the detection
unit,
directly
illuminates
the C/W light,
but activates
only the MASTER
ALARM switch lights
of the master alarm circuit.
One event signal,
from MSFYNstations
through
the UDL portion
of
"CRRW ALERT," originates
the communications
system.
This system status light
can only be
extinguished
by a second signal
originating
from the MSFN.
The master alarm circuit
alerts
crew-members whenever abnormal
conditions
are detected.
This is accomplished
visually
by illumination
of remote MASTER ALARM switch-lights
on MDC-1, MDC-3, and LEB-122.
An
audio alarm tone, sent to the three headsets,
aurally
alerts
the crew.
The output signal
of the tone generator
is a square wave that is
alternately
750 and 2000 cps, modulated
at 2.5 times per second.
Although
the tone is audible
above the conversation
level,
it does not render
normal conversation
indistinct
or garbled.
When the crew has noted the
abnormal condition,
the master alarm lights
and the tone generator
are
deactivated
and reset by depressing
any one of the three MASTER ALARM
switch-lights.
This action
leaves the systems status
lights
illuminated
and resets the master alarm circuit
for alerting
the crew if another
abnormal condition
should occur.
The individual
systems status
lights
will
remain illuminated
until
the malfunction
or out-of-tolerance
or the NORMAL-BOOST-ACK switch
(MIX!-3) is
condition
is corrected,
positioned
to ACK.
The C!&WSpower supplies
include
sensing and switching
circuitry
that insure
unit self-protection
should high-input
current,
or high- or
low-output
voltage
occur.
Any of these fault
conditions
will
cause the
illumination
of the master alarm lights
and the C/W system status
light.
The tone generator,
however, will
not be activated
because it requires
the 12 V dc output from the malfunctioned
power supply for its operation.
The crew must manually
select
the redundant
power supply to return
the
C&WSto operation.
This is accomplished
by repositioning
the CAUTION/
WARNING-POWERswitch on MDC-2. In so doing, the C/W status
light
is
extinguished,
but the master alarm circuit
remains activated,
requiring
it to be reset.
Incorporated
into the C&WSis the capability
to test the lamps of
Position
1 of the CAUTION/
systems status
and master alarm lights.
WARNING-LAMP TEST switch tests the illumination
of the left-hand
group
of status
lights
on MDC-2 and the MASTER ALARM switch-light
on MDC-1.
A-9
__
,..-
..__
_--._,__-,..--l.--=/~~-
.I
I
"I.
--
-""."
.~
_-__-_-..,
4
--1__1
~^
~---.---l--.---....-..I
.__I-
___-___(,__
Position
2
hand group
located
on
on LEB-122
tests
the MASTER ALARM switch-light
on MDC-3 and the rightMASTER ALARM light,
of status
lights
on MDC-2. The third
LEB-122, is tested by placing
the CONDITION LAMPS switch
to TEST.
The position
of the CAUTION/WARNING - CSM-CM switch (MIX-2)
Before CM-SM separation,
establishes
the systems to be monitored.
systems in both the CM and SM are monitored
for malfunction
or out-oftolerance
conditions
with this switch in the CSM position.
Positioning
the switch to CM deactivates
systems status
lights
and event indicators
associated
with SM systems.
(MDC-2) permits
The CAUTION/WARNING - NORMAL-BOOST-ACK switch
For most
variable
modes of status
and alarm light
illumination.
of the mission,
the switch is set to the NORMAL position
to give
of abnormal condition
normal C&WSoperation;
that is, upon receipt
signals,
all systems status
lights
and master alarm lights
are
During
the
ascent
phase,
the
switch
is set to
capable of illumination.
the BOOST position,
which prevents
the MASTER ALARM switch-light
on
This prevents
possible
confusion
on MDC-1
MDC-1 from illuminating.
between the red MASTER ALARM light
and the adjacent
red ABORT light.
The ACK switch position
is selected
when the crew desires
to adapt
or if a continuously
illuminated
systems status
their
eyes to darkness,
will
activate
While in this mode, incoming signals
light
is undesirable.
To determine
the
only the master alarm lights
and the tone generator.
the crew must depress either
MASTER ALARM switchabnormal condition,
This illuminates
the applicable
systems status
light
on MDC-1 or -3.
The systems
and deactivates
and resets the master alarm circuit.
light,
status
light
will
remain illuminated
as long as the switch-light
is
as long as the abnormal
depressed.
However, it may be recalled
condition
exists
by again pressing
either
switch-light.
A stowable
tone booster
is added to the caution
to allow all three astronauts
to sleep simultaneously
Stowage of this unit during non-use periods
removed.
A3.
and warning system
with the headsets
is under locker
The unit consists
of a power plug, tone booster,
and a photosensitive
device which can be used on the left
or right
side of the
The power connection
is made to the UTILITY receptacle
command module.
which provides
an audible
signal,
on MDC-15 or 16. The tone booster,
is mounted by velcro pad to the left-hand
or right-hand
girth
shelf.
The photo-sensitive
device is mounted by velcro
over the MDC-1 or
MDC-3 MASTER ALAP&J lamp.
Since the MASTER ALARM is triggered
activate
the
monitored
symptom, it will
A-95
by any caution/warning
tone booster
until
the
MASTER ALARM is extinguished
by a manual reset.
In the event of a
caution/warning
system power supply failure,
this unit will
provide
the audio alarm.
Electrical
power distribution
.- The C&WSreceives
power from
the MNA & MNB buses (see fig.
A2.10-1).
Two circuit
breakers,
located
on MDC-5, provide
circuit
protection.
Closure of either
circuit
breaker will
allow normal system operation.
Figure
A2.10-l.-
C&WSpower
distribution
diagram.
Operational
Limitations
and Restrictions
With the CAUTION/WARNING - NORMAL-BOOST-ACK switch in the
BOOST position
during ascent,
the MASTER ALARM switch-light
on
MDC-1 will
not illuminate
should a malfunction
occur.
The master alarm
circuit
reset capability
of the light
is also disabled
during this
This requires
the MASTER ALARM switch-light
on MDC-3 to be
time.
used exclusively
for monitoring
and resetting
functions
during boost.
Several peculiarities
should be noted in regard to the CAUTION/WARNING POWERswitch.
Whenever this switch is moved from or through the OFF
position
to either
power supply position,
the master alarm circuit
is
activated,
requiring
it be reset.
Also, switching
from one power supply
to another
(when there is no power supply failure)
may cause the C/W
system status
light
to flicker
as the switch passes through the OFF
position.
Should both power supplies
fail,
the C&WSis degraded to the extent
that the complete master alarm circuit,
as well as those system status
lights
that illuminate
as the result
of analog-type
input signals,
are
rendered inoperative.
This leaves only those status
lights
operative
that require
event-type
input signals.
They include
the following
SM
and CM lights:
CMC, ISS, BMAG 1 TEMP, BMAG 2 TEMP, SPS ROUGHECO,
PITCH GMBL 1, PITCH GMBL 2, YAW CMBL 1, YAW GMBL 2, 02 FLOW HI,
FC BUS DISCONNECT., AC BUS 1, AC BUS 1 OVERLOAD, AC BUS 2, AC BUS 2
OVERLOAD, MN BUS A UNDERVOLT, MN BUS B UNDERVOLT, and CREWALERT. The
C/W light
will
be operative
only while the CAUTION/WARNING - POWER
switch is in position
1 or 2.
The CAUTION/WARNING - CSM-CM switch must be in the CSM position
in order to conduct a lamp test of those system status
lights
associated
with SM systems.
The status lights
of CM systems may be tested with the
switch in either
position.
Circuit
design permits
a complete lamp
test to be conducted with the CAUTION/WARNING switch in the NORMAL or
In the BOOST position,
all lamps except the MASTER
ACK position
only.
ALARM light
on MDC-1 may be tested.
Normally,
each abnormal condition
signal
will
activate
the C&WS
and illuminate
the applicable
master alarm circuit
and tone generator,
However, after
initial
activation
of any status
systems status
light.
light
that monitors
several
parameters,
and reset of the MASTER ALARM,
any additional
out-of-tolerance
condition
or malfunction
associated
with
the same system status
light
will
not activate
the MASTER ALARM until
the first
condition
has been corrected,
thus extinguishing
the status
light.
A-97
__=I
.
..“.-_
_..“...
,.,,
“‘
..--.
-~-___
--
-----^1
Each crewmember's audio control
panel has a power switch which
will
allow or inhibit
the tone signal
from entering
his headset.
The
AUDIO-TONE position
allows the signalto
pass on to the headset,
while the AUDIO position
inhibits
the signal.
PART A2.11
MISCELLANEOUS SYSTEMS DATA
Introduction
Miscellaneous
systems data pertain
in other systems.
These items consist
(G-meter),
and uprighting
system.
to items
of timers,
that are not covered
accelerometers
Timers
Two mission timers
(electrical)
and two event timers
(electrical/
mechanical)
are provided
for the crew in the command module.
One
mission
timer is located
on panel 2 of the MDC and the other on panel
306 in the left-hand
forward equipment bay.
Each mission timer has
provisions
for manually
setting
the readout
(hours,
minutes,
and
seconds),
and the capability
of starting,
stopping,
and resetting
to
The numerical
elements are electroluminescent
lamps and the
zero.
intensity
is controlled
by the NUMBRICS light
control
on panels MDC-8
and LEB-100.
The event timers
are located
on MDC-1 and -306 in the
left-hand
forward equipment bay, and provide
the crew with a means of
monitoring
and timing
events.
Ali timers reset and start
automatically
when lift-off
occurs,
and the timer located
on MDC-1 will
be automatically
reset and restarted
if an abort occurs.
The event timers
are integrally
illuminated
by an internal
electroluminescent
lamp and controlled
by the
INTEGRAL light
controls
located
on MDC-8 and LEB-100.
Accelerometer
(G-meter)
The accelerometer
or G-meter @DC-l) provides
the crew with a
visual
indication
of spacecraft
positive
and negative
G-loads.
This
meter is illuminated
by an internal
electroluminescent
lamp and controlled
by the INTEGRAL light
control
on MDC-8.
Command Module
The CM uprighting
system
the CM has assumed a stable,
consists
of three inflatable
valves,
two air compressors,
able bags are located
in the
The
in the aft compartment.
Uprighting
System
is manually
controlled
and operated
after
The system
inverted
floating
attitude.
air bags, two relays,
three solenoid-control
control
switches,
and air lines.
The inflat'34 forward compartment and the air compressors
control
switches
and circuit
breakers
are
located
in the crew compartment.
The switches
control
relays which
are powered by the postlanding
bus and the relays
control
power to the
compressors which are powered by battery
buses A and B.
(See figure
A2.11-1.)
EATTLRV
BUS B
FLTIPL
BUS
CT
c
xLA
>
EPS BAT
BUSA
IRHLE-2N
)'
FLOAT BAG 1
BAT A
IMDC-8)
I
5A
Y
25A
UPRIGHT svs
CWPR
1
EAT A
IAHEB-ml
2CA
I )BUSBEPS BAT
2%
WEB-2??I
I
5A
)
P
FLOAT BAG
L WC-81
-----
------
1
/
I
i
Q,
FLOAT BAG R
BAT B
iMDC~81
-----
1
I
I
VfNl
I
!
_--m---e
NORMCLOSED
CONTRNVALVL
M. l-V BAG
I
I
Pizmid---_---e-m-
MRM
CIOSED
,----m-m
COMROLVALVf
NO.Z+VBAG
CONTROL VALVE
W.3+ZBAG
4
T
52
Figure
A2.11-l.-
Sequential
systems
operational/functional
r
1
COMPRtSSOR COlvTROt
MOTOR SWITCHLS
IRHLB-2981
diagram.
FLOAT BAG 1L switch controls
inflation
Functional
description.switch 2R controls
inflation
of the air
of the air bag on -Y axis,
and switch 3 CTR controls
inflation
of the air bag
bag on the +Y axis,
Two of the bags are
on the +Z axis of the CM (see fig.
A2.11-1).
If the
45 inches in diameter;
the other bag is 24 inches in diameter.
the crewmember at station
1 initiates
CM becomes inverted
after
landing,
filling
of the three bags by setting
the FLOAT BAG lL, 2R, and 3 CTR
When the CM is uprighted,
the three FLOAT BAG switches
switches
to FILL.
will
be set to OFF. A 4.25f0.25-psi
relief
valve is located
in the inlet
Backup relief
valves set at 13.5 psi are located
in the
of each bag.
outlet
of each compressor.
-.
A-102
___l.s^_.--..-" .--..^..._I
-.-.
"."._
FART A2.12
CREWPERSONAL EQUIPMENT
This section
contains
the description
and operation
of Contractorand NASA-furnished
crew personal
equipment and miscellaneous
stowed
equipment that is not described
in other sections
of the handbook.
All
mador items are identified
as Contractor-furnished
equipment
(CFE) or
Government-furnished
(NASA) property
(GFP - synonymous with GFE).
The crew equipment
usage in SM2A-03-BLOCK.
a.
Intravehicular
(a)
(b)
(c)
(d)
(e)
(2)
Spacesuit
Assembly
Biomedical
Harness and Belt
Constant Wear Garment (CWG)
Flight
Coveralls
Pressure Garment Assembly (PGA)
Associated
Umbilicals,
Adapters,
Extravehicular
(a)
(b)
(c)
C.
of operational
Spacesuits
(1)
b.
is presented
in the general
order
A brief
outline
is as follows:
Spacesuit
Assembly
Liquid-Cooled
Garment (LCG)
PGA with Integrated
Thermal Meteroid
Associated
Equipment
G-Load Restraints
(1)
Crewman Restraint
Harness
(2)
Interior
and Straps
(3)
Hand Bar
Zero-g
and Equipment
Handhold
Restraints
(1)
Rest Stations
(2)
Velcro
(3)
Straps
and Snap Restraint
A-103
Areas
Garment
(ITMG)
d.
Internal
Sighting
(1)
Window Shades
(2)
Mirrors
(3)
Crewman Optical
(4)
LM Active
(5)
Window Markings
(6)
Miscellaneous
External
e.
f.
and Illumination
Alignment
Docking
Sighting
(2)
Running
(3)
EVA Floodlight
(4)
EVA Handles
with
(5)
Rendezvous
Beacon
Lights
(1)
Flight
(2)
Inflight
(3)
Cameras
(4)
Accessories
(1)
Water
(2)
Food
RL Disks
Aids
Data File
Toolset
& Miscellaneous
Waste Bags
Pilot's
Preference
Kits
Fire Extinguishers
Oxygen Masks
Utility
Outlets
Scientific
Instrumentation
Support
A-104
(...._.__--. .-_.-_.-^ -------.-
Aids
Spotlight
Operational
63. Crew Life
(COAS)
Target
and Illumination
Exterior
(a)
(b)
(c)
(d)
(e)
(f)
Sight
Aids
(1)
Mission
Aids
(PPKs)
Outlets
(3)
The Galley
(4)
Waste Management
(5)
Personal
h.
Medical
1.
Radiation
ii.
Postlanding
k.
System
Hygiene
Supplies
and Equipment
Monitoring
Recovery
(1)
Postlanding
(2)
Swimmer Umbilical
(3)
Recovery
Beacon
(4)
Snagging
Line
(5)
Seawater
Pump
(6)
Survival
Kit
Equipment
System
and Measuring
Aids
Ventilation
Stowage
Equipment
Ducts
and Dye Marker
PART AZ.13
DOCKING AND TRANSFER
Introduction
This section
identifies
system and the operations
Docking operational
tions
and text describe
docking.
These activities
the physical
associated
with
characteristics
of the
docking and separation.
docking
sequence .- The following
sequence of illustrathe general
functions
that are performed
during
will
vary with the different
docking modes.
After the spacecraft
and third
stage have orbited
the earth,
possibly
up to three revolutions,
the third
stage is reignited
to place the
spacecraft
on a translunar
flight.
Shortly
after translunar
injection,
the spacecraft
transposition
and
docking phase takes place (fig.
A2.13-1).
When the CSM is separated
from the third
stage, docking is achieved by maneuvering
the CSM close
enough so that the extended probe (accomplished
during earth orbit)
engages with the drogue in the LM. When the probe engages the drogue
with the use of the capture latches,
the probe retract
system is activated
to pull the LM and CSM together.
Upon retraction,
the LM tunnel
ring will
activate
the 12 automatic
docking ring latches
on the CM and effect
a pressure
seal between the
modules through
the two seals in the CM docking ring face.
After the
two vehicles
are docked, the pressure
in the tunnel
is equalized
from
the CM through
a pressure
equalization
valve.
The CM forward hatch is
removed and the actuation
of all 12 latches
is verified.
Any latches
not automatically
actuated will
be cocked and latched manually by the
crewman.
The LM to CM electrical
umbilicals
are retrieved
from their
stowage position
in the LM tunnel
and connected to their
respective
connectors
in the CM docking ring.
The vehicle
umbilicals
supply the power to release the L&J from the
SLA. Once the hold-down
straps are severed,
four large springs
located
at each attachment
point push the two vehicles
apart,
and the combined
CSM/LM continues
towards the moon.
A-106
-
H
A-107
. _..-_
_. _.
..~- - ..._”..--_ _.._.1-- I-__~
_-__..^_.__
~,l_____~
__---
~-s_--.--
Once in lunar orbit,
the tunnel
is repressurized.
The probe assembly and drogue assembly are removed from the tunnel
and stowed in the
CM. The pressure
in the LM is equalized
through
the LM hatch valve,
With the pressure
equalized,
the IM hatch is opened and locked in the
open position
to provide
a passageway between the two modules.
After two crewmen transfer
to the LM, the CM crewman retrieves
the
drogue from its stowage location
in the CM, passes it through the
tunnel,
and helps to install
and lock it in the tunnel.
The drogue may
be installed
and locked by the LM crewmen, if they choose.
The probe
assembly is then retrieved
from its stowage location
in the CM and
installed
and preloaded
to take all the load between the modules.
This
accomplished,
the LM hatch is closed by the I.&l crewmen.
The 12 docking
latches
are released
and cocked by the crew-man in the CM so that the
latches
are ready for the next docking operation.
The CM forward hatch
is reinstalled
and checked to assure a tight
seal.
The modules are now
prepared
for separation.
The probe EXTEND RELEASE/RETRACT switch in the CM (MDC-2) is placed
energizing
the probe extend latch.
The probe
in the EXTEND position,
extends and during extension
will
activate
a switch energizing
an enternal electrical
motor to unlock the capture latches.
After the probe
extends,
the I.M pulls
awsy from the CM and descends to the lunar surface.
If the switch is not held until
the probe reaches f'ull extension,
the
The
capture latches
will
reengage to hold the two vehicles
together.
switch would then have to be reactivated
and separation
performed with
the RCS.
it will
be several
hours before the first
man steps
After landing,
few hours are spent checking the IM ascent
foot on the moon. The first
This completed,
the cabin is depressurized
and one
stage and resting.
Following
a short period,
of the crewmen descends to the lunar surface.
Lunar surface
activities
the se'cond crewman
descends to the surface.
will
vary for each mission.
engine
Following
is fired
completion
using the
of the lunar surface
exploration
the ascent
depleted
descent stage as a launch platform,
After rendezvous
and docking in lunar orbit,
the LM crewmen transfer back to the CM, After the CSM and I&i pressures
have equalized,
the
LM crew opens the LM hatch while the CM pilot
removes the tunnel hatch.
The drogue and probe are removed and stowed in the LM. Lunar samples,
film,
and equipment to be returned
to earth are transferred
from the LM
to the CM. Equipment in the CM that is no longer needed is put into the
LM, and the LM hatch is closed,
the CM hatch is replaced,
and the seal
checked.
A-108
The IN is then released
by firing
the separation
system (detonating
cord) located
around the circumference
of the docking ring,
thus serving
the ring and abandoning
the I&I (fig.
A2.13-1).
This completed,
the CM
SPS engine is fired,
placing
the spacecraft
in a return
trajectory
toward
the earth.
Functional
Description
The docking system is a means of connecting
and disconnecting
the
LM/CSM during a mission
and is removable to provide
for intravehiclular
transfer
between the CSM and L&l of the flight
crew and transferrable
equipment.
The crew transfer
tunnel,
or CSM/LM interlock
area, is a passageway
between the CM forward bulkhead
and the LM upper hatch.
The hatch
relationship
with the docking hardware is shown in fig.
AZ.l3-2.
(The
figure
does not show the installed
positions.)
For descriptive
purposes
that portion
of the interlock
area above the CM forward bulkhead to the
docking interface
surface
is referred
to as the CM tunnel.
That portion
of the interlock
outboard
of the L&I upper hatch extending
to the docking
interface
surface
is referred
to as the I&l tunnel.
The CM tunnel
incorporates
the CM forward hatch, probe assembly,
docking ring and seals,
and
the docking automatic
latches.
The L&I tunnel
contains
a hinged pressure
drogue locking
mechanism,
drogue support
fittings,
drogue assembly,
hatch,
and L&f/CM electrical
umbilicals.
A-109
A-
-.-- .
..-
--
. ..--.
--.
.-.
.._ _... “... /..
-__-
-.“.”
110
,,..--,
-.~,
.__ 1-1--1-
-
-~-~---~
PART A3
LUNAR MODULE SYSTEMS DESCRIPTION
INTRODUCTION
This part includes
descriptions
of the LM, the LM - spacecraft-tolunar module adapter
(SLA) - S-IVB connections,
the LM-CSM interfaces,
and LM stowage provisions
are included
in this
chapter.
These data were
extracted
from the technical
manual LMA 790-3-I&I,
Apollo Operations
Handbook, Lunar Module, Volume 1, dated February
1, 1970.
LM CONFIGURATION
The LM (fig.
A3-1) is designed for manned lunar landing
missions.
It consists
of an ascent stage and a descent stage; the stages are
joined together
at four interstage
fittings
by explosive
nuts and bolts.
Subsystem continuity
between the stages is accomplished
by separable
interstage
umbilicals
and hardline
connections.
Both stages function
as a single
unit during lunar orbit,
until
separation
is required.
Stage separation
is accomplished
by explosively
severing
the four interstage
nuts and bolts,
the interstage
umbilicals,
and the water lines.
All other hardlines
are disconnected
automatically
at stage separation.
The ascent stage can function
as a single
unit to
accomplish
rendezvous
and docking with the CSM. The overall
dimensions
of the LM are given in figure
A3-2.
Station
reference
measurements
(fig.
Al-l)
ar e established
as follows:
a. The Z- and Y-axis
station
reference
measurements (inches)
at a point where both axes intersect
the X-axis
at the vehicle
vertical
centerline:
the Z-axis extends forward and aft of the
intersection;
the Y-axis,
left
and right.
The point
of intersection
is established
as zero.
start
b.
The +Y-axis
measurements
increase
-Y-axis measurements increase to the left.
-Z-axis
measurements
increase
forward
(+Z)
to the
right
Similarly,
and aft
(-Z)
from zero;
the
the tZ- and
from zero.
station
reference
measurements
(inches)
start
at a
C. The X-axis
This reference
design reference
point identified
as station
+X200.000.
point is approximately
128 inches above the bottom surface
of the
footpads
(with the landing
gear extended);
therefore,
all X-axis
station
reference
measurements are +X-measurements.
A-111
-. .,",..
__.
_---. -___
. ,.-.____,,.-__._.
--ll~--ll-".-----*l
__Xsll_---_~-.
~-^_-l_-_--ll--_
Figure
A3-l.-
LM configuration.
::- 112
(--14,
I”-=
7
28’ a 2”
19’ 10.65”
t
Y 8.86”
1
Figure
A3-2.-
LM overall
A-113
dimensions.
d
FORWARD
.-
Ascent
Stage
The ascent stage, the control
center and manned portion
of the
IN, accommodates two astronauts.
It comprises three main sections:
the crew compartment,
midsection,
and af't equipment bay.
The crew
compartment
and midsection
make up the cabin, which has an approximate
overall
volume of 235 cybic feet.
The cabin is climate-controlled,
Areas other than the cabin are
and pressurized
to 4.8 - 0 .2 psig.
unpressurized.
Crew Compartment.
- The crew compartment is the frontal
area of
the ascent stage; it is 92 inches in diameter
and 42 inches deep.
This
is the flight
station
area; it has control
and display
panels,
armrests,
body restraints,
landing
aids, two front windows, a docking window,
and an alignment
optical
telescope
(AOT).
Flight
station
centerlines
are 44 inches apart;
each astronaut
has a set of controllers
and armcontrol,
and display
panels are along the
Circuit
breaker,
rests.
upper sides of the compartment.
Crew provision
storage space is
The main control
and display
panels are canted
beneath these panels.
and centered between the astronauts
to permit sharing
and easy scanning.
between the flight
stations,
is used in
An optical
alignment
station,
conjunction
with the AOT. A portable
life
support system (PLSS) donning
slightly
aft of the optical
alignstation
is also in the center aisle,
ment station.
The crew compartment has 12 control
panels:
Control
and display
two main display
panels (1 and 2)
and display
panels (fig.
A3-3):
two
center
panels
(3 and 4) that
that are canted forward 10 degrees,
slope down and aft 45 degrees towards the horizontal,
two bottom side
panels (5 and 6), two lower side panels (8 and 12), one center side panel
rate display
(14), two upper side panels (11 and 161, and the orbital
(ORDEAL) panel aft of panel 8.
earth and lunar
A-114
-- ._-. -
___I__ -."111..-- -_--
---. -
.-I_^_
Figure
A3-3.-
Cabin
interior
(looking
forward).
Panels 1 and 2 are located
on each side of the front
face assembly
Each panel is constructed
of two 0.015-inchi
centerline,
at eye level.
thick
aluminum-alloy
face sheets,
spaced 2 inches apart by formed
channels.
The spacer channels
are located
along the sheet edges;
additional
channels,
inboard
of the edge channels,
reinforce
the sheets.
This forms a rigid
box-like
construction
with a favorable
strength-toweight ration
and a relatively
high natural
frequency.
Four shock mounts
support
each panel on the structure.
Panel instruments
are mounted to
the back surface
of the bottom and/or to the top sheet of the panel.
The instruments
protrude
through
the top sheet of the panel.
All dial
faces are nearly
flush with the forward face of the panel.
Panel1
contains
warning lights,
flight
indicators
and controls,
and propellant
quantity
indicators.
Panel 2 contains
caution
lights,
flight
indicators
and controls,
and Reaction
Control
Subsystem (RCS) and Environmental
Control
Subsystem (ECS) indicators
and controls.
Panel 3 is immediately
below panels 1 and 2 and spans the width
of these two panels.
Panel 3 contains
the radar antenna temperature
indicators
and engine,
radar,
spacecraft
stability,
event timer,
RCS
and lighting
controls.
Panel 4 is centered between the flight
stations
and below panel 3.
Panel 4 contains
attitude
controller
assembly (ACA) and thrust
translation
controller
assembly (TTCA) controls,
navigation
system indicators,
and LM guidance
computer (I&C) indicators
and controls.
Panels 1 through
4 are within
easy reach and scan of both astronauts.
Panels 5 and 6 are in front
of the flight
stations
at astronaut
waist height.
Panel 5 contains
lighting
and mission
timer controls,
engine start
and stop pushbuttons,
and the X-translation
pushbutton.
Panel 6 contains
abort guidance controls.
The panel is
Panel 8 is at the left
of the Commander's station,
canted up 15 degrees from the horizontal;
it contains
controls
and
displays
for explosive
devices,
audio controls,
and the TV camera
connection.
Panel 11, directly
above panel 8, has five angled surfaces
that
contain
circuit
breakers.
Each row of circuit
breakers
is canted
15 degrees to the line of sight so that the white band on the circuit
breakers
can be seen when they open.
A-116
canted
tions,
Panel 12 is at the right
of the LM Pilot's
station.
The panel is
up 15 degrees from the horizontal;
it contains
audio, communicaand communications
antennas controls
and displays.
Panel 14, directly
above panel 12, is canted
It contains
controls
and displays
the horizontal.
distribution
and monitoring.
up 36.5 degrees
for electrical
from
power
Panel 16, directly
above panel 14, has four angled surfaces
that
Each row of circuit
breakers
is canted
contain
circuit
breakers.
15 degrees to the line of sight so that the white band on the circuit
breakers
can be seen when they open.
(ORDEAL) panel is
The orbital
rate display
- earth and lunar
immediately
aft of the panel 8. It contains
the controls
for obtaining
I,M attitude,
with respect
to a local horizontal,
from the LCC.
Windows:
Two triangular
windows in the front
face assembly provide
visibility
during descent,
ascent,
and rendezvous
and docking phases of
the mission.
Both windows have approximately
2 square feet of viewing
area; they are canted down to the side to permit adequate peripheral
and downward visibility.
A third
(docking)
window is in the curved
overhead portion
of the crew compartment shell,
directly
above the
Commander's flight
station.
This window provides
visibility
for
docking maneuvers.
All three windows consist
of two separated
panes,
vented to space environment.
The outer pane is made of Vycor glass with
a thermal
(multilayer
blue-red)
coating
on the outboard surface
and an
antireflective
coating
on the inboard
surface.
The inner pane is made
of structural
glass.
It is sealed with a Race seal (the docking
window inner pane has a dual seal) and has a defog coating
on the
outboard
surface
and an antireflective
coating
on the inboard surface.
Both panes are bolted
to the window frame through
retainers.
All three windows are electrically
heated to prevent
fogging.
The heaters
for the Commander's front window and the docking window
receive
their
power from 115-volt
ac bus A and the Commander's 28-volt
dcbus,
respectively.
The heater for the LM Pilot's
front
window receives
power from 115-volt
ac bus B. The heater power for the Commander's front
window and the docking window is routed through the AC BUS A: CDR WIND
HTR and HEATERS: DOCK WINDOWcircuit
breakers,
respectively;
for the
LM Pilot's
front window, through the AC BUS B: SE WIND HTR circuit
breaker.
These are 2-ampere circuit
breakers
on panel 11. The temperature of the windows is not monitored
with an indicator;
proper heater
operation
directly
affects
crew visibility
and is, therefore,
visually
determined
by the astronauts.
When condensation
or frost
appears on a
window, that window heater is turned on.
It is turned off when the
abnormal condition
disappears.
When a window shade is closed,
that
window heater must be off.
The midsection
structure
(fig.
A3-4) is a ring-stiffened
Midsection.The bulkheads
consist
of aluminum-alloy,
chemically
semimonocoque shell.
milled
skin with fusion-welded
longerons
and machined stiffeners.
The
midsection
shell
is mechanically
fastened
to flanges
on the major
structural
bulkheads
at stations
+'Z27,OO and -Z27.00.
The crew compartment shell
is mechanically
secured to an outboard
flange of the +Z27.00
+x294.643
and +X233.500,
The up-per and lower decks, at stations
bulkhead.
are made of aluminum-alloy,
integrally
stiffened
and
respectively,
The lower deck provides
structural
support
for the ascent
machined.
The upper deck provides
structural
support
for the docking
stage engine.
tunnel
and the overhead hatch.
Figure
A3-A.-
Cabin
interior
(looking
aft).
Two main beams running
fore and aft,
integral
with those above
the crew compartment,
are secured to the upper deck of the midsection;
they support the deck at the outboard
end of the docking tunnel.
The
aft ends of the beams are fastened
to the aft bulkhead
(-Z27.000)
which has provisions
for bolting
the tubular
truss members that s&port
both aft interstage
fittings.
Ascent stage stress
loads applied
to the
front beam are transmitted
through
the two beams on the upper deck to
the aft bulkhead.
Descent
Stage
The descent stage is the unmanned portion
of the LM. It contains
the descent engine propellant
system, auxiliary
equipment for the
end scientific
experiment
packages to be placed on the
astronauts,
The descent stage structure
provides
attachment
and
lunar surface.
support points
for securing
the LN within
the spacecraft-lunar
module
adapter
(SLA).
LM - SLA - S-IVB
Connections
At earth launch,
the LM is within
the SLA, which is connected to
the S-IVB booster.
The SLA has an upper section
and a lower section.
to which the landing
gear is attached,
provide
attachment
The outriggers,
The LM is mounted
points
for securing
the IN to the SLA lower section.
to the SLA support
structure
on adjustable
spherical
seats at the apex
it is held in place by a tension
holddown
of each of the four outriggers;
Before the IN is removed, the upper
strap at each mounting point.
These
section
of the SLA is explosively
separated
into four segments.
which
are
hinged
to
the
lower
section,
fold
back
and
are
then
segments,
The IN is then explosively
forced away from the SLA by spring thrusters.
released
from the lower section.
LM-CSM Interfaces
A ring at the top of the ascent stage provides
for joining
the LM to the CSM. The ring,
which is
clamping mechanisms in the CM, provides
structural
drogue portion
of the docking mechanism is secured
The drogue is required
during docking operations
to
See figure
A3-5 for orientation
CM-mounted probe.
CSM.
a structural
compatible
continuity.
below this
mate with
of the LM
interface
with the
The
ring.
the
to the
The crew transfer
tunnel
(LPI-CM interlock
Crew transfer
tunnel.area) is the passageway created between the I&I overhead hatch and the
CM forward pressure
hatch when the I&l and the CSM are docked.
The
A-119
..- .--. * _- -.--- --l_--_“.“.l..-
I... .-_--.-__ll-___ll_“_~
tunnel
permits
intervehicular
exposure to space environment.
transfer
of crew and equipment
without
Final
docking latches:
Twelve latches
are spaced equally
about
periphery
of the CM docking ring.
They are placed around and within
CM tunnel
so that they do not interfere
with probe operation.
When
the latches
insure
structural
continuity
and pressurization
secured,
between the LM and the cT4, and seal the tunnel
interface.
Umbilical:
An electrical
umbilical,
in the
tunnel,
is connecCed by an astronaut
to the CM.
be made without
drogue removal.
CREWMAN
OPTICAL
ALIGNMENT
SIGHT
STANDOFF
;CSM-ACTIVE
ALIGNMENT
\
CROSS
DOCKING
TARGET)
LM
IX portion
of the
This connection
can
LM -Y-AXIS
ACQUISITION
CH
LM COAS LINE
OF SIGHT POST
PITCHOVER POSITION
CSM
AND
\
ACOUISITION
ORIENTATION
LIGHT (TIP)
Figure
A3-5.-
STANDOFF
CROSS ANU
ALIGNMENT
STRIPS
(LM -ACTIVE
DOCKlnv.I.-.
ALIGNMENT
TARCEll
LM-CSM reference
A-120
axes.
the
the
Docking hatches.the CSM has a single,
is not removable.
It
The IN has a single
docking
(overhead)
hatch;
integral,
forward hatch.
The LM overhead hatch
is hinged to open 75 degrees into the cabin.
The drogue assembly
Docking drogue.provisions
for mounting in the LM portion
The drogue may be removed from either
end
and may be temporarily
stowed in the CM or
Propulsion
System (SPS) burns.
One of the
a locking
mechanism to secure the installed
is a conical
structure
with
of the crew transfer
tunnel.
of the crew transfer
tunnel
the LM, during Service
three tunnel
mounts contains
drogue in the tunnel.
Docking probe.The docking probe provides
initial
CM-LM coupling
and attenuates
impact energy imposed by vehicle
contact.
The docking
probe assembly consists
of a central
body, probe head, capture latches,
pitch
arms, tension
linkages,
shock attenuators,
a support
structure,
probe stowage mechanism, probe extension
mechanism, probe retraction
system, an extension
latch,
a preload
torque shaft,
probe electrical
umbilicals,
and electrical
circuitry.
The assembly may be folded
for
removal and stowage from either
end of the transfer
tunnel.
The probe head is self-centering.
When it centers
in the drogue
the three capture latches
automatically
engage the drogue socket.
The
capture latches
can be released 'oy a release handle on the CM side of
the probe or by depressing
a probe head release button
from the I&l
side, using a special
tool stowed on the right
side stowage area inside
the cabin.
Visual
alignment
aids are used for final
alignment
Docking aids.of the IN and CSM, before the probe head of the CM makes contact with
The LM +Z-axis will
align
50 to 70 degrees from the
the drogue.
CSM -Z-axis
and 30 degrees from the CSM +Y-axis.
The CSM position
represents
a 180-degree pitchover
and a counterclockwise
roll
of
60 degrees from the launch vehicle
alignment
configuration.
An alignment
target
is recessed into the LM so as not to protrude
into the launch configuration
clearance
envelope or beyond the LM envelope.
-~46.30O and -20.203,
has a
The target,
at approximately
stations
radioluminescent
black standoff
cross having green radioluminescent
disks on it and a circular
target
base painted
fluorescent
white with
black orientation
indicators.
The base is 17.68 inches in diameter.
Cross members on the standoff
cross will
be aligned
with the orientation
indicators
and centered within
the target
circle
when viewed at the
intercept
parallel
to the X-axis
and perpendicular
to the Y-axis
and
Z-axis.
A-121
Stowage Provisions
The IM has provisions
for stowing
crew personal
equipment.
The
equipment includes
such items as the docking drogue; navigational
star charts and an orbital
map; umbilicals;
a low-micron
antibacteria
filter
for attachment
to the cabin relief
and dump valve;
a crewman's
medical kit;
an extravehicular
visor
assembly (EWA) for each astronaut;
a special
multipurpose
wrench (tool B); spare batteries
for the PLSS
packs; and other items.
-_
A-122
PART A4
MISSION CONTROL CENTER ACTIVITIES
INTRODUCTION
The Mission
Control
Center (MCC) is located
at the Manned Spacecraft
Center in Houston, Texas.
The MCC contains
the communications,
computer
display
and command systems to effectively
monitor
and control
the
Apollo spacecraft.
These data were extracted
from information
furnished
by Flight
Operations
Directorate,
Manned Spacecraft
Center.
Flight
operations
are controlled
from the MCC. The MCC contains
two flight
control
rooms, but only one control
room is used per mission.
a Mission
Operations
Control
Room (MOCR),
Each control
room, called
is capable of controlling
individual
Staff Support Rooms (SSR) located
adjacent
to the MOCR. Both the MOCR's and the SSR's operate on a
To accomplish
this,
the various
flight
control
functions
24-hour basis.
Figures A4-1 and A4-2
and consoles
are staffed
by three g-hour shifts.
show the floor
plans and locations
of personnel
and consoles in the
Figure Ah-3 shows MOCR activity
during the Apollo
MOCR and the SSR's.
and figure
Ah-4 shows the MOCR and SSR organizational
13 flight,
structure.
9. Flight Surgeon: Directs all operational medical activities concerned with the mission, including
the status of the flight crew.
10. Spacecraft Communicator: Voice communications with the astronauts, exchanging information
on the progress of the mission with them.
11. Flight Dynamics Officer: Monitors and
evaluates the flight parameters required to achieve
a successful orbital flight; gives “GO” or “ABORT”
recommendations to the Flight Director.
12. Retrofire Officer: Monitors impact prediction displays and is responsible for determination
of retrofire times.
14. Booster Systems Engineer: Monitors pro
pellant tank pressurization systems and advises the
Bight crew and/or Flight Director of systems abnormalities.
15. Guidance Officer: Detects Stage I and
Stage II slowrate deviations and other programmed
events, verifies proper performance of the Inertial
Guidance System, commands onboard computation
function and recommends action to the Flight
Director.
16. Network Controller: Hasdetailedoperational control of the Ground Operational Support System
network.
17. DepartmentofDefenseRepresentative:
Overall control of Department of Defense forces supporting
the mission, including direction of the deployment
of recovery forces, the operation of the recovery
communications network, and the search, location
and retrieval of the crew and spacecraft.
1. Flight Operations Director: Responsible for
successful completion of mission flight operations
for all missions being supported.
2. Mission Director: Overall mission responsibility and control of flight test operations, which
include launch preparation. In Project Mercury there
were no alternative mission objectives that could
be exercised other than early termination of the
mission. The Apollo missions, however, offer many
possible alternatives which have to be decided in
real time.
3. Public Affairs Officer: Responsible for providing information on the mission status to the
public.
4. Flight Director: Responsible for detailed
control of the mission from lift-off until conclusion
of the flight.
5. Assistant Flight Director: Responsible tothe
Flight Director for detailed control of the mission
from lift-off through conclusion of the flight; assumes
the duties of the Flight Director during his absence.
6. Experiments and Flight Planning: Plansand
monitors accomplishment of flight planning and
scientific experiment activities.
7. Operations and Procedures Officer: Responsible to the Flight Director for the detailed implementation of the MCC/Ground Operational Support
Systems mission control procedures.
8. Vehicle Systems Engineers: Monitor and
evaluate the performance of all electrical, mechanical and life support equipment aboard the spacecraft (this includes the Agena during rendezvous
missions).
Figure
Ah-l.-
Personnel
and console
A-124
-__ ~. _ _---_l----_l_-
--
locations.
Legend of symbols
0
Clock
i
Zone fire alarm panel
lg
“A”
power panel
[Bi
“6”
power panel
8
Frre extinguisher
[XI Air duct
m
Ozone exhaust
Room no.
Room name
310
311
312
312A
313
314
316
319
Flight dynamtcs SSR
Vehicle systems SSR
Life systems SSR
Flight crew SSR
Operation and procedures SSR
AiSEP
SSR
Control and display terminal
Display and timing
324A
Meteaological
324
3248 1
327
Recovery
::;A
329
330
zz:“,
332
Figure
Ab-2.-
Floor
center
control
Recovery control display projection
Simulation control
Summary display projection
Mission operations control room no. 1
f Comm booth
Visitors
viewing
plan of MOCKand SSR's.
area
Fliyht
1
I
I
I
Network
CCWllOlli?l
Spacecraft
Commll,~Icator
AD
Act~wt~fs
DfflCU
I
IST
Flight
Dynalu~cs
officer
I
I
I
l
Retrofire
Officer
Guidance
Dff1cer
/ Yaw
;
CSM Systeiw
Engineers
EECDM
: GNC
LM Systems
Engineers
TELCDM
! CONTROL
Booster Systeins
EMU
E”iJlWXS
’
-J
1
t
Figure
A4-4.-
NOCR and SSR organizational
structure.
MISSION OPERATIONS CONTROL ROOM
The MOCR was the center for mission
control
operations.
The prime
controlpositions
were stationed
in this area.
The MOCRwas broken
Responsibilities
of the groups were
down into three operations
groups.
as follows :
a.
Mission
(I)
Command and Control
Group
Mission
Director
(MD)
The MD was responsible
for
overall
conduct
(2) Flight
Operations
Director
(FOD)
The FOD was responsible
for the interface
Director
and management.
of the mission.
between
the Flight
(3) Flight
Director
(FD)
The FD was responsible
for MOCR decisions
and actions
concerning
vehicle
systems, vehicle
dynamics,
and MCC/MSFN
operations.
(4) Assistant
Flight
Director
(AFD)
The AFD was responsible
for assisting
the
in the performance
of his assigned duties.
Flight
Director
(5) Flight
Activities
Officer
(FAO)
The FAO was responsible
for developing
the flight
plan.
and coordinating
(6) Department
of Defense Representative
(DOD)
The DOD Representative
was responsible
for coordination
and direction
of all DOD mission
support forces and sites.
(7) Assistant
DOD Representative
The Assistant
DOD Representative
was responsible
for
ing the DOD Representative
in the performance
of his
assisttask.
(8) Network
Controller
(NC) (NETWORK)
The Network Controller
was responsible
Director
for the detailed
operational
analysis
of the MSFN.
to the Flight
control
and failure
(9) Assistant
Network Controller
The Assistant
Network Controller
assisted
the Network
Controller
in the performance
of his duties and was responsible
for all MCC equipment and its ability
to support.
~-128
-.__
-_“~
i_
.,l----.l..
-.-.._-_
.
.._-
-
.-..-
-.~_l_-
l____.-_ll_l__~-,-“,~-.~..l_--
-.-
(10)
Public
Affairs
Officer
(PAO)
The PA0 was responsible
for keeping
on the progress
of the mission.
the public
informed
(11)
Surgeon
The Flight
Surgeon was responsible
to the Flight
Director
for the analysis
and evaluation
of all medical activities
concerned with the flight.
(12)
Spacecraft
Communicator
(CAPCOM)
The Spacecraft
Communicator was responsible
to the Flight
Director
for all voice communications
with the flight
The CAPCOMalso served in conjunction
with FAO as
crew.
This position
was manned by
a crew procedures
advisor.
a member of the backup flight
crew.
(13) Experiments
Officer
(EO) (EXPO)
The primary
function
of the EO was to provide
overall
operational
coordination
and control
for the Apollo Lunar Surface
Experiment
Package (ALSEP), and the Lunar Geology ExperiThe coordination
was with the various
MOCR
ment (LGE).
operational
positions
and the ALSEP SSR; the Principal
Investigators,
Management, the Program Officer,
Goddard,
The EO was also
and the Manned Space Flight
Network.
responsible
to the Flight
Director
for providing
ALSEP
and LGE status
and any ALSEP or LGE activities
that could
have an effect
on the Apollo mission.
b.
Systems
Operations
Group
(MOCR)
(1)
Environmental,
Electrical,
and Communications
(EECOM)
The CSM EECOM Engineer was responsible
to the FD for
monitoring
and troubleshooting
the CSM environmental,
and sequential
systems.
electrical,
(2)
Guidance,
Navigation,
and Control
(GNC)
The GNC Engineer was responsible
to the Flight
Director
for monitoring
and troubleshooting
the CSM guidance,
control,
and propulsion
systems.
navigation,
(3) TELCOM
The LM Environmental
and Electrical
Engineer was responsible to the FD for monitoring
and troubleshooting
the
electrical,
and sequential
systems.
LM environmental,
A-129
._
. ."- -"
_..-. ...._..----
~I.*ll _--. l.--".-
_--
.*----1.--1".1_1_..
(4) CONTROL
and Control
Engineer was
The LM Guidance,
Navigation,
responsible
to the Flight
Director
for monitoring
and
troubleshooting
the LM guidance,
navigation,
control,
and propulsion
systems.
(5) Booster
Systems Engineer
(BSE)
The Booster Systems Engineers'
responsibilities
delegated
as follows:
(a)
(6)
C.
BSE 1 had overall
responsibility
for the launch veIn addition,
hicle
including
command capability.
BSE 1 was responsible
for all S-IC and S-II stage
functions.
(b) BSE 2 had prime responsibility
functions
with the exception
for all S-IVB
of command.
(c)
for all
exception
BSE 3 had prime responsibility
unit
(IU) functions
with the
stage
instrument
of command.
Apollo
Communications
Engineer
(ACE) (INCO) and Operations
and Procedures
Officer
(O&P) (PROCEDURES)
The INCO and O&P shared a console and responsibility.
The INCO's prime responsibility
to the Flight
Director
was for monitoring
and troubleshooting
the CSM, LM, TV,
PLSS, and erectable
antenna communication
systems.
He
was also responsible
for execution
of all commands assoThe O&P's prime
ciated with the communication
systems.
responsibility
to the Flight
Director
was for the detailed
implementation
of the MCC/MSFN/GSFC/KSC mission
control
The O&P was also responsible
for
interface
procedures.
scheduling
and directing
all telemetry
and DSE voice
He also developed all communication
inputs
playbacks.
and changes to the ground support timeline.
Flight
(1)
were
Dynamics
.-
Group
Flight
Dynamics Officer
(FIDO)
The Flight
Dynamics Officer
participated
in prelaunch
checkout designed to insure
system readiness,
monitored
powered flight
events and trajectories
from the standpoint of mission
feasibility;
monitored
reentry
events
and updated impact point estimates
as
and trajectories,
required.
A-130
._-
",__,_
^ _,._,._
- _._____________
_-_l_l._ ,_--,
*-II.-L.I"._._
".---.--+--
..---_ - _,........._ __ ,__~..,.
- _;..II~
._.. .------
...- ^_._
--III
(2)
Retrofire
Officer
(REPRO)
The Retrofire
Officer
participated
in prelaunch
checkout
designed to insure
system readiness
and maintained
an
updated reentry
plan throughout
the mission.
(3) Guidance
Officer
(GUIDO) and YAW
The Guidance Officer
participated
in prelaunch
checkout
designed to insure
system readiness
and performed
the
guidance monitor
functions
during power flight
and spacecraft
initialization.
The GUIDO was also responsible
for
CSM and LM display
keyboards
(DSKY) as well as CMC and LGC
The second Guidance Officer
(YAW) had
command updates.
the same duties
except that he was not responsible
for
command functions.
MCC SUPPORT ROOMS
Each MOCR group had a staff
support room (SSR) to support all activities
required
by each MOCR position.
These SSR's were strategically
located
in areas surrounding
the MOCR's and were manned by the various
personnel
of a given activity.
a.
Staff
Support
Room
(1) Flight
Dynamics SSR
The Flight
Dynamics SSR was responsible
to the Flight
Dynamics Group in the MOCR for providing
detailed
analysis
of launch and reentry
parameters,
maneuver requirements,
It also, with the assistance
of
and orbital
trajectories.
the Mission Planning
and Analysis
Division
(WAD), provided
real-time
support
in the areas of trajectory
and guidance
to the MOCR Flight
Dynamics team on trajectory
and guidance
An additional
service
required
provided
interface
matters.
between the MOCR Flight
Dynamics team and parties
normally
outside
the Flight
Control
team such as Program Office
spacecraft
contractor
representatives,
representatives,
et cetera.
(2)
Flight
Director's
SSR
The Flight
Director's
SSR was responsible
for staff
support
to the Flight
Director,
AFD, Data Management Officer.,
and
to the Apollo CommuniFAO. This SSR was also responsible
cations
Engineer
in the MOCR for monitoring
the detailed
The SSR was also
status of the communication
systems.
Ground Timeline
responsible
for two TV channel displays:
and Flight
Plan.
A-131
(3) Vehicle
Systems SSR
The Vehicle
Systems SSR was responsible
to the Systems
Operations
Group in the MOCR for monitoring
the detailed
status
and trends of the flight
systems; avoiding,
correcting, and circumventing
vehicle
equipment failures;
and
detecting
and isolating
vehicle
malfunctions.
After the
S-IVB was deactivated,
the portable
life
support
system
engineer
and the Experiments
Officer
occupied the two
booster
consoles in the Vehicle
Systems SSF.
(4)
Life Systems SSR
The Life Systems SSR was responsible
to the Life Sy-stems
Officer
for providing
detailed
monitoring
of the physiological
and environmental
data from the spacecraft
concerning the flight
crew and their
environment.
(5)
Spaceflight
Meteorological
Room
The Spaceflight
Meteorological
Room was responsible
Mission
Command and Control
Group for meteorological
space radiation
information.
(6)
to the
and
Space Environment
Console (SEC) (RADIATION)
The Space Environment
Console was manned jointly
by a Space
Environment
Officer
(SEO) from the Flight
Control
Division
and a Space Environment
Specialist
from the Space Physics
Division.
During mission
support,
the SE0 was responsible
the proper operation
of the confor the console position,
of all necessary
activities
and
sole, and the completion
The SEC was the central
collecting
and coordiprocedures.
nating
point
at MSC for space radiation
environment
data
during mission periods.
(7) Spacecraft
Planning
and Analysis
(SPAN) Room
The SPAN Room was the liaison
interface
between the MOCR,
the data analysis
team, vehicle
manufacturers,
and KSC
Launch Operations.
During countdown and real-time
operations,
the SPAN team leader initiated
the appropriate
action
necessary
for the analysis
of spacecraft
anomalies.
(8) Recovery
Operations
Control
Room (ROCR)
The Recovery Operations
Control
Room was responsible
for
the recovery
phase of the mission
and for keeping the Flight
Director
informed
of the current
status
of the recovery
Additionally,
the Recovery Operations Control
operations.
Room provided
an interface
between the DOD Representative
and the recovery
forces.
A-132
.^ -
.-.-_-.-
. __..-__
_.-_
__-.v,-___
_..--~.-.--- .._....
.~--l~
-..-.,.-l__ll_l__l__l__ll_l(".~."--".~------
(9) ALSEP SSR
The ALSEP SSR was responsible
to the Experiments
Officer,
Lunar Surface Program Office,
and Principal
Investigators
for providing
detailed
monitoring
of ALSEP central
station
and experiments
data.
for all
The SSR was also responsible
scheduling
of activities,
commanding, and data distribution
to appropriate
users.
MISSION SUPPORT AREAS
The two primary
support areas for the MOCR flight
control
team were
the CCATS area and the RTCC area located
on the first
floor
of the MCC.
These two areas of support and their
operational
positions
interfaced
with the MOCR flight
control
team.
Communications,
Command, and Telemetry
System
(CCATS)
The CCATS was the interface
between the MCC and MSFN sites.
CCATS was a hardware/software
configuration
(Univac 494 computer) having
the capability
to provide
for the reception,
transmission,
routing,
processing,
display
and control
of incoming,
outgoing,
and internally
generated
data in the areas of telemetry,
command, tracking,
and administrative
information.
The CCATS consoles were augmented with various
high-speed
printers
(HSP) and TTY receive-only
(RO) printers
adjacent
the CCATS operational
organiFigure Ah-5 illustrates
to the consoles.
CCATS personnel
interfaced
with the MOCR flight
control
team
zation.
were as follows:
a.
Command Support Console
This console was a three-position
support element whose operators were concerned with the total
command data flow from the generation
and transfer
of command loads from the RTCC to the verification
of space
The three comvehicle
acceptance
following
uplink
command execution.
mand positions
were:
(1) Real-Time
(2)
Command Controller
Command Load Controller
(3) CCATS Command Controller
(RTC)
(LOAD CONTROL)
(CCATS CMD)
A-133
. ..-."I____".x.
..-..-..
..*,...,II, ,_ t.__,.._
- .---
-.--l.l---
Telemetry
Instrumentation
Control
Console
This console was a two-position
support element whose operators
were concerned with the telemetry
control
of incoming
data from the MSFN.
telemetry
program control
was exercised
on the incoming
data.
Certain
The two telemetry
positiozis
were:
b.
(1)
Telemetry
Instrumentation.
(2)
CCATS Telemetry
Controller
Controller
(TIC)
(CCATS TM)
Instrumentation
Tracking
Controller
Console
This console was a two-position
support element whose operators
were concerned with the tracking
radar support involving
the spacecraft
and ground systems operations
and configurations.
The two tracking
positions
were:
C.
(1)
Instrumentation
Tracking
Controller
(TRK)
(2) USB Controller
a.
Central
Processor
Control
Console
This console was a two-position
support element and provided
the facilities
for monitoring
and controlling
selected
software
and
hardware functions
applicable
to the configuration
of the CCATS computer
The two positions
were:
complex.
(1)
Central
Processor
Controller
(CPC)
(2)
Central
Processor
Maintenance
and Operations
Communications
Controller
Console
of this
console provided
The operators
management between MCC and MSFN elements.
(M&O)
e.
Real-Time
Computer
Complex
overall
communications
(RTCC)
The RTCC provided
the data processing
support for the MCC. It
storage
and limit
sensing,
traaccomplished
the telemetry
processing,
command load generation,
display
jectory
and ephemeris calculations,
and many other necessary
logic processing
and calculations.
generation,
The RTCC supported
both MOCR's and as such had two divisions
known as
each capable of supporting
one MOCR.
computer controller
complexes,
known as the misEach complex was supported
by two IBM 360 computers,
sion
operations
computer
(MOC) and the dynamic standby computer (DSC).
the RTCC
The DSC served as backup to the MOC. Figure Ah-6 illustrates
A brief
description
of the
operational
organization
for each complex.
RTCC positions
f'~llo~~.
A-134
._".
_
-..-_.-.-.I^--.-_-I_-. ."--., -, ...---fl"..."Li ",._-II .._^
.."-.-l-.-_-."--‘.~l-.--.--
Flujilt
D,rector
Network
Colltroller
Figure
A&?.-
WATS operational
organization.
Flight
Director
+
Flight Dynamics
Processing
Controller
Tracking Data
Selection
Controller
Figure
Telemetry
Processing
Controller
~4-6.- RTCC operational
organization.
Network and Command
Processing
Controller
a.
RTCC Director
Controlled
and coordinated
the
activities
of the
two computer
complexes.
b.
RTCC Computer Supervisor
(Computer Sup)
Responsible
for the operational
control
of the
complex.
Tracking
Data Selection
Controller
(Data Select)
Monitored
the tracking
data being processed
in the RTCC and
insured
the data used as input to the MOCR and SSR displays
was the
Evaluated
the quality
of tracking
data received
during
best obtainable.
Evaluated
the trajecthe launch phase and selected
the source of data.
tory determinations
and was responsible
for the various
related
displays.
Informed the MOCR Flight
Dynamics Officer
concerning
the quality
and
status
of the data.
c.
Flight
Dynamics Processing
COntrOller
(COIIIPUter
WnmiCS)
Controlled
and monitored
all trajectory
computing requirements
requested
by MOCR flight
dynamics personnel
and MOCR recovery
activities.
Performed evaluation
and analysis
of the predicted
trajectory
quantities
as they related
to the mission plan.
d.
Network and Command Processing
Controller
(Computer Command)
Coordinated
with MOCR personnel
who had command responsibility
and transfer
of requested
command
and directed
the generation,
review,
loads.
e.
Telemetry
Processing
Controller
(Computer TM)
This position
had access to all telemetry
data entering
and
leaving
the RTCC and interfaced
with the MOCR and SSR positions
using
Duties included
monitoring
telemetry
input data, coortelemetry
data.
monitoring
computer generated
telemetry
disdinating
input requests,
and
keeping
the
MOCR
aware
of
the
telemetry
processing
status.
plays,
f.
NOTE
From ALSEP deployment
to splashdown TRK and
TIC will
be re.sponsible
for scheduling
sites
This will
to support the scientific
package.
include
calling
up of sites and data/command
handling
to MCC.
A-137
I .. . .
_,..
-.I---...
,.__-_._..-I _..I".-.- -.-11..~.
.- _I-.^
..__ _^^..-.-_l-"____
This
page left
blank
intentionally.
A-138
(1.
..-.,-
._..~.
I---
..--
~--..-
-..-
_--__l__l-
.-l-t_l----,---
.
.
-
“.
_
^...
.
_.
.
.._
-
_.-l-l.-
.-__.-
-
.I
-_,,.-
-_.
PART A5
EXCERPTSFROMAPOLLO FUEL CELL AND
CRYOGENICGAS STORAGESYSTEMFLIGHT SUPPORTHANDBOOK
The information
contained in this part was extracted
from the
Apollo Fuel Cell and Cryogenic Gas Storage System Flight Support Handbook,
and Power
dated February 18, 1970. It was prepared by the Propulsion
Division
of the Manned Spacecraft
Center.
The text was taken from SecSection 3.0 Cryogenic Gas
tion 2.0 Fuel Cell Operation and Performance,
Storage System Operation and Performance, Section 4.0 Instrumentation
Subsystem Maland Caution and Warning, Section 5.0 Fuel Cell/Cryogenic
Section 7.0 Fuel Cell/Cryogenic
Subsystem Hardware
function
Procedures,
Description.
2.0
FUEL CELL OPERATION AND PERFORMANCE
The fuel cell operation
and performance
are described
by nominal
system performance
and operational
data for both ground and flight
environments
and fuel cell response to a variety
of component malfunctions.
Nominal system performance
and operational
data are presented
in
curve and table format to assist
in rapid reference.
The data
include
procedures
and curves for making rough estimates
of radiator
performance.
Apollo 10 and 11 flight
data were used to generate a
portion
of the curves used for evaluating
radiator
performance.
Fuel cell response of measured parameters (temperature,
voltage,
etc.)
to specific
component malfunctions
make up the remainder of the data
The curves are adequately
noted to allow application
presented.
without
written
procedures.
The fuel cell operation
and performance data assist
the user in
evaluating
fuel cell performance,
identification
of flight
anomalies
and provide a basis for developing
corrective'actions.
The sources of the data were the original'NASA
Apollo Elock II Fuel
Cell, Cryogenic Gas Storage System,and Flight
Batteries
Flight
Support
Handbook','dated
September 1968, NASA-MSC, North American Rockwell,
Pratt and Whitney, Beech Aircraft
and Boeing-Houston.
These data were
reviewed and found to be accurate as of December 1969.
A-140
..--
-..“...
,------mm
.__--._l”-.-.--l_l..-
_I
..,,..
-ill_--.-
-__.....
_ “.-.-----ill-.-
--.,__I
.-----
2.1
FUEL CELL SYSTEM OPERATIONAL PARAMETERSSUMMARY(Continued)
NOMINAL FUEL CELL PRESSURIZED SYSTEM VOLUMES
Hydrogen
250 in3
Loop
88 in3
Oxygen Loop
Nitrogen
Loop
3098 in3
Glycol
Loop (Fuel
Glycol
Accumulator
Net Fuel Cell
(Fuel
Glycol
Average NR Glycol
Volume
Estimated
107 in3
Cell)
Cell)
30 in3
Volume
Plumbing
and Radiator
Glycol
Loop Volume
Fuel Cell
117 in3
66 in3
(20 in3 water
-glycol
removed
from accumulator)
183 in3 = 0.79 gallons
TOTAL SYSTEM SPEC LEAKAGE INTO BAY IV
Hydrogen System,
Cells (3)
Oxygen System,
5.3 x 10-j
Fuel Cell
(3) nitrogen
and Fuel
scc/sec
of Helium
system
1.6 x lO-4 scc/sec
of Helium
SYSTEM PRESSURESUMMARY
SUPPLY PRESSURES
SUPPLY PRESSURES
(PSIA)
REGULATEDPRESSURES
S/N 650769 AND ON
DEAD
ABOVE
ABSOLUTE
NITROGEN
BAND
PSIA
PRESSURE
PSI
PSI
NOMINAL
MINIMUM
Hydrogen
245 + 15
100
57.90 - 67.60
6.20 - 11.35
.2-.4
Oxygen
900 + 35
150
57.90 - 67.95
6.20 - 11.7
.5-.?
Nitrogen
1500
165
50.20 - 57.75
---
2.02.15
DEL IVERY PRESSURE
Water 62 psia
PRESSURELIMITS
Maximum water system discharge
back pressure
Maximum reactant
vent back pressure 16 psia
A-142
59.55 psia
2.1
FUEL CELL SYSTEM OPERATIONAL PARAMETERSSUMMARY(Continued)
ENVIRONMENTAL CONTROLSYSTEM WATER SYSTEM PRESSURES
Potable
Water Tank
Water Relief
25 psia
Valve
5.5 psid
Water Tank Vent Valve
44 psia
Cabin
6 . 0 +'24
-.
Relief
+ 2, Plus
Valve
cabin
pressure
?r 1
+ 4
FUEL CELL GROUNDHEATER POWERSETTINGS
STARTUP HEAT SCHEDULE
1
ZONE
1
AMPERES
1
NORMALOPERATION HEAT SCHEDULE
1 ZONE 1
SEA LEVEL OPERATION
1
VACUUMOPERATION
1
1.2 - 1.6 amperes
0 amperes
2
8.0 - 12.0 amperes
0 amperes
3
1
1.2 - 1.6 amperes
0 amperes
I
DRYOUT HEAT SCHEDULE
ZONE
SEA LEVEL OPERATION
1
1.75 - 2.05 amperes
2
As
required
to maintain
460°F
to 485'F
skin
VACUUMOPERATION
1.5 - 1.65 amperes
21.0 - 22.5 amperes
temperature.
Approximately 23.9 amps.
3
1.75 - 2.05 amperes
A-143
1.5 - 1.65 amperes
1
1
2.1
FUEL CELL SYSTEM OPERATIONAL PARAMETERSSUMMARY(Continued)
FUEL CELL DISCONNECT OVERLOADDATA
OVERLOADCURRENT DATA
Required
Disconnect
Delay (set)
100 minimum
25 - 300
8- 150
2-8
:, 622 - 1.2
0:42 - 0.76
0.24 - 0.55
Test
Delay
(se4
Transfer
Time
(set)
No transfer
0.046
0.046
0.046
i"8
5.81
0.776
1.07
0.572
0.470
0.046
0.046
0.046
FUEL CELL DISCONNECT REVERSE CURRENTDATA
REVERSE CURRENT DATA
Transfer
Time
(set)
Required
Disconnect
Delay (set)
Test
Delay
(set)
4
20
No trip
i - 10
2.10
0.046
:i
1 - 1.3
1.11
1.22
0.046
Load
':%j
No transfer
2.1
FUEL CELL SYSTEM OPERATIONAL PARAMETERSSUMMARY(Continued)
REACTANT CONSUMPTIONAND WATER PRODUCTION
LOAD
AMps
O2 lb/hr
H2°
H2 lb/hr
lb/hrs
cc/hr
0.5
0.0102
.001285
.01149
5.21
1
0.0204
.002570
.02297
10.42
2
0.0408
.005140
.04594
20.84
3
0.0612
.007710
.06891
31.26
4
0.0816
.010280
.09188
41.68
5
0.1020
.012850
.11485
52.10
6
0.1224
.015420
.13782
62.52
7
0.1428
.017990
.l6079
72.94
8
0.1632
.020560
.18376
83.36
9
0.1836
.023130
.20673
93.78
10
0.2040
.025700
.2297
104.20
15
0.3060
.038550
.34455
156.30
20
0.4080
.051400
.45940
208.40
25
0.5100
.064250
.57425
260.50
30
0.6120
.077100
.68910
312.60
35
0.7140
.089950
.80395
364.70
40
0.8160
.10280
.91880
416.80
45
0.9180
.11565
1.03365
468.90
50
1.0200
.I2850
1.1485
521.00
55
1.1220
.A4135
1.26335
573.10
60
I .2240
.I5420
1.3782
625.20
65
1.3260
.16705
1.49305
677.30
70
1.4280
.I7990
I .6079
729.40
75
1.5300
.I9275
1.72275
781.50
80
1.6320
.20560
1.83760
833.60
85
1.7340
.218L,5
1.95245
885.70
90
1.8360
.23130
2.06730
937.90
95
1.9380
.24415
2.18215
989.90
100
2.0400
.25700
2.2970
1042.00
FOFMJLAS:
O2 = 2.04
x 10 -2 I
H20 = IO.42
H2 = 2.57
x 1O-3 I
H20 =
A-145
2.297
cc/Amp
Hr
x 10 -2 lb/Amp
Hr
3.0
CRYOGENIC GAS STORAGESYSTEM OPERATION AND PERFORMANCE
The cryogenic
system operation
and performance
are described
by
nominal system performance
and operational
data for both ground
and flight
environments.
Nominal system performance
and operational
data are presented in
curve and table format to assist
in rapid reference.
The curves,
with the exception
of those used for heat leaks and pressure change
are adequately
noted to allow application
without
written
rates,
The data include
formulas,
methods, and curves for
procedures.
calculating
cryogenic
tank heat leak s and pressure change rates for
both equilibrium
and non-equilibrium
(stratified)
conditions.
Apollo 7 and 8 flight
data were used to provide a comparison of
equilibrium
(calculated)
tank pressure cycle time to actual flight
pressure cycle time for a variety
of tank quantities.
and performance data assist
the user in
The fuel cell operation
evaluating
cryogenic
system performance,
identification
of flight
and provide
a basis for developing
corrective
actions.
anomalies,
The sources of the data were the original'NASA
Apollo Block II Fuel
Cell, Cryogenic Gas Storage System, and Flight
Batteries
Flight
Support Handbook',' dated September 196S, NASA-MSC, North American
Rockwell,
Pratt and Whitney, Beech Aircraft
and Boeing-Houston.
These data were reviewed and found to be accurate as of December 1969.
3.1
CRYOGENIC SYSTEM OPERATIONAL PARAMETERSSUMMARY
Hydrogen
Oxygen
TANK WEIGHT (PER TANK)
Empty (Approx.)
80.00
lb.
90.82
lb.
Usable
28.15
lb.
323.45
lb.
29.31
lb.
330.1
lb.
Fluid
Stored Fluid
indication)
(100%
4%
Residual
Maximum Fill
Quantity
TANK VOLUME (PER TANK)
2%
30.03
lb.
6.80
FT3
337.9
lb.
4.75 FT3
TANK FLOW RATE (PER TANK)
Max. for
Max. for l/2
Relief
Valve
4.03
1.02 lbs/hr
10 Minutes
lbs/hr
10.40 lbs/hr
hour
Max Flow
6 lbs/hr
130°F
@
26 lbs/hr
130°F
@
TANK PRESSURIZATION
Heaters (2 elements
Flight
Resistance
per tank)
Maximum Voltage
Power
Total Heater
Heat Input Per Tank
(2 Elements)
Ground
Resistance
Maximum Voltage
Power
Total Heater
Heat Input Per Tank
(2 Elements)
* Conversion
Factor:
1 watt
78.4 ohms per
element
28 V DC
10 watts per
element*
10.12 ohms per
element
28 V DC
77.5 watts per
element*
68.2 BTU/Hr
528.6 BTU/Hr
78.4 ohms per
element
65.0 V DC
10.12 ohms per
element
65.0 V DC
54.0 watts
element*
417.5 watts
element*
368 BTU/Hr
= 3.41 BTU/Hr
per
2848 BTU/Hr
per
3.1
CRYOGENIC SYSTEM OPERATIONAL PARAMETERSSUMMARY
Hydrogen
Oxygen
260 psia
225 psia
10 psia
935 psia
865 psia
30 psia
115/200 V
400 cps
3.5 watts per
motor*
23.0 BTU/Hr
115/200 V
400 cps
26.4 watts per
motor*
180 BTU/Hr
Normal Operating
245 515 psia
900 +35 psia
Spec Min. Dead Band of
Pressure Switches
10 psi
30 psi
Pressure Switch
Open Pressure
Close Pressure
Deadband
idax.
Min.
Min.
Destratification
Motors
Motors Per Tank)
Voltage
(2
Power - Average
Total Average Motor
Heat Input Per Tank
SYSTEEl PRESSURES
Relief
Valve
Note:
Relief Valves are Referenced to Environmental Pressure,
therefore
Pressure at
Sea Level (psig) will
be same value in
vacuum (psia)
273 psig
983 psig
::::
285 psig
1010 psig
Min.
268 psig
965 psig
+ 10
90 _ 2. psid
75 2 7.5 psid
-425 to 80°F
-3OO'F to 80°F
N.A. for 113
and Subs.
N.A. for 114
and Subs.
Crack
Full
Flow
Reseat
Outer
Burst
Tank Shell
Disc
Nominal
Burst
Pressure
SYSTEM TEMPERATURES
Stored
Fluid
Heater Thermostat
Temp. Protection)
(Over
Open
Max.
80°F + 10
80°F + 10
Close
t.1i n .
-2OO'F
-75'F
* Conversion
Factor:
1 watt
= 3.41 BTU/Hr
3.1
CRYOGENIC SYSTEM OPERATIONAL PARAMETERSSUMMARY
Hydrogen
Oxygen
7.25 BTU/HR
(.0725 #/hr)
27.7 BTU/HR
(.79 #/hr)
400 see H2/HR/
Valve
400 see 02/HR/
Valve
0.736 x 1O-6 lbs
H2/HR/Valve
9.2 x 1O-6 lbs
02/HR/Valve
600 HRS @
Cryogenic Temps
and operating
pressure -225
psia
600 HRS @
Cryogenic Temps.
and operating
pressure -865
psia
TANK HEAT LEAK (SPEC PER TANK)
Operating
(dQ/dM @ 14O'F)
VALVE MODULE LEAKAGE RATES
External
LIFE
A-149
,--._ I,, ..- .._-. .__.,
. .~._.~ .... _-..I._.-...-__._ ..""
I. _ ,, _.......-_..
_.~.-- ___.
-x
1._._-,._.
__,,_.__. ....__,"__.__--.l
. _ ,...-
4.0
INSTRUMENTATION AND CAUTION AND WARNING
The tabular
data presented
in Tables 4.1 and 4.2 list
instrumentation
measurements and specify
instrumentation
range, accuracy and bit
value,
if applicable.
All of the data in Tables 4.1 and 4.2 can be
used for system monitoring
during ground checkout.
Table 4.1 lists
data displayed
to the crew and telemetered
from the vehicle
to the
Manned Space Flight
Network (MSFN) during missions.
Table 4.2 lists
data available
only for system monitoring
during ground checkout.
Event indications
displayed
to crew during flight
are noted in
Table 4.2.
The instrumentation
sensor location,
with the exception
of voltage
and current
data, can be found by referring
to the fuel celllcryogenie schematics
located
in Section 7.0.
Voltage and current
readout
and schematic locations
can be obtained
by referring
to North
American Rockwell drawings V37-700001, Systems Instrumentation,
and
V34-900101, Integrated
System Schematics Apollo CSM, respectively.
The Caution and Warning System monitors
the most critical
fuel cell/
cryogenic
measurements and alerts
the flight
crew to abnormal system
The data presented
in Table 4.1 are specification
nominal
operation.
caution
and warning limits
for the applicable
measurements.
Malfunctions procedures,
Section 5.0, are provided
for problem isolation
as
a result
of a caution
and warning alarm.
The source of the data was North American Rockwell Measurement
Systems End-to-End Calibrated
Accuracy Tolerances,
TDR68-079,
dated January 10, 1969 and the original
Flight
Support Handbook.
A-150
__.--
.
_
_
.
..-..
-..
^_
. ..---_--
.-_”
-..,^
i*_,r
._“_
,__-
“.-,-.----f.ell
_...
_Ili”---.---L-
__l_-l__l---I_LLll,-_-___.l---,.--------~.
“-I.---
TABLE 4.1
INSTRUMENTATION/CAUTION AND WARNINGSUMMARY
EAStiREMENT
NUMBER *
MEASUREMENT
NAME
CCO206V
DC Bus Voltage
A
CCO207V
DC Bus Voltage
B
SC2113C
FC 1 Current
SC2114C
FC 2 Current
sc2115c
FC 3 Current
SC206OP
FC 1 N2 Press
SC2061P
FC 2 N2 Press
SC2062P
FC 3 N2 Press
SC2066P
FC 1 02 Press
SC2067P
FC 2 02 Press
SC2068P
FC 3 02 Press
SC2069P
FC 1 H2 Press
SC207OP
FC 2 H2 Press
~~2071~
FC 3 i-12 Press
t See note,
page 4-11
RANGE
T
At JRACY
PERCENT
ACTUAL
T T
BIT
VALUE
CAUTION AND
WARNING
SET NGS
HIGH
LOW
26.25
-
o-45 volts
20.94
iO.42V
0.178
O-100 amps
11.07
11.07 a
0.395
O-75 psia
-14.30
~3.22 psia
0.295
-
O-75 psia
i4.30
23.22 psia
0.295
-
-
O-75 psia
14.30
k3.22 osia
0.295
-
-
1
TABLE 4.1
INSTRUMENTATION/CAUTION AND WARNINGSUMMARY (Continued)
IEASUREMENT
NUMBER *
MEASUREMENT
NAME
SC2081T
FC 1 Cond Ex Temp
SC2082T
FC 2 Cond Ex Temp
SC2083T
FC 3 Cond Ex Temp
SC2084T
FC 1 Skin Temp
SC2085T
FC 2 Skin Temp
SC2086T
FC 3 Skin Temp
SC2087T
FC 1 Rad Out Temp
SC2088T
FC 2 Rad Out Temp
SC2089T
FC 3 Rad Out Temp
SC2090T
FC 1 Rad In Temp
SC2091T
FC 2 Rad In Temp
SC2092T
FC 3 Rad In Temp
SC2139R
FC 1 H2 Flow Rate
SC2140R
FC 2 H2 Flow Rate
SC2141R
FC 3 H2 Flow Rate
* See nnte,
page 4-11
RANGE
T
AC RACY
ACTUAL
PERCENT
BIT
VALUE
CAUTION AND
WAF NG
SE1 NGS
HIGH
im-
145-250°F
~2.18
t2.29'F
0.417
150°F
175'F
80-550°F
21.15
i5.40°F
1.94
360°F
5OO'F
-50 to +300°F
i . 71
5.98'F
1.38
-3OOf
-50 to +300°F
k1.71
?5.98'F
1.38
O-O.2 lb/hr
+_lO.O
kO.020 lb/hr
0.00079
0.0
0.16
1
TP E 4.1
TRUMENTATION/CA UN AND W-SUMMARY
EASUREMENT
NUMBER *
MEASUREMENT
NAME
SC2142R
FC 1 O2 Flow Rate
SC2143R
FC 2 02 Flow Rate
SC2144R
FC 3 O2 Flow Rate
SCOO3OQ
H2 Tank 1 Qty
SCOO31Q
H2 Tank 2 Qty
SCOO32Q
O2 Tank 1 Qty
SCOO33Q
O2 Tank 2 Qty
SCOO37P
O2 Tank 1 Press
SCOO38P
O2 Tank 2 Press
SCOO39P
H2 Tank 1 Press
SCOO4OP
H2 Tank 2 Press
SC0041T
O2 Tank 1 Temp
SC0042T
02 Tank 2 Temp
SC0043T
H2 Tank 1 Temp
SC0044T
H, Tank 2 Temp
* See note,
page 4-11
RANGE
O-l.6
lb/hr
AC RACY
PERCENT
ACTUAL
dO.0
kO.160 lb/hr
lConti
l-
BIT
VALUE
0.0063
1
YEi-
ION AND
WARNING
SF
0.0
1.27
o- 100%
+2.68
2.68%
0.4%
-
-
o- 100%
52.68
2.68%
0.4%
-
-
k26.8 psia
4.23
800
950
kg.38 psia
1.48
220
270
50-1050
psia
O-350 psia
22.68
+2.68
-320 to +~O'F
k2.68
+10.85'F
1.57
-
-
-420 to -2OO'F
+2.68
t6.03'F
0.867
-
-
TABLE 4.1
Il~STRtiMENTATIOid/CAUTION AND WARNINGSUM?lARY (Continued)
1EASUREMEHT
NUMBER*
MEASUREMENT
NAME
RANGE
}7TmP
KY
ACTUAL
CAUTIOH AND
WARNING
SE TINGS
Low
HIGH
BIT
VALUE
SC216OX
FCl pH High
Normal - High
Event
SC2161X
FC2 pH High
Normal - High
Event
-
SC2162X
FC3 pH High
Normal - High
Event
-
*SC0050Q
H2 Tank 3 Qty
O-100%
'*SCOO51Q
02 Tank 3 Qty
O-100%
,*SCOO52P
H2 Tank 3 Press.
O-350 psia
'*scoo53P
02 Tank 3 Press.
50-1050 psia
* See note, page A-160
** CSM 112 through 115 only
,
,I
TABLE 4.2
GROUNDTEST INSTRUMENTATION
MEASUREMENT
NAME
BIT
VALUE
RANGE
I
FC 1 Bus A
FC 2 Bus A
FC 3 Bus A
SC2125X **
FC 1 Bus B
SC2126X **
FC 2 Bus B
SC2127X **
FC 3 Bus B
scoo92x
Pressure
O2 tanks
scoo93x
Motor Switch
Close O2 tanks
low
1 & 2
Pressure
H2 tanks
scoo95x
Motor Switch
Close H2 Tanks l&2
SCO36OV
Fan Notor Oper
Tank 1 O2
SCO361V
Fan blotor Oper
Tank 2 O2
page A-160
Event
Open - Close
Event
Normal - Low
Event
Open - Close
Event
l&2
scoo94x
* See note,
Normal - Low
low
1 & 2
** Data also
displayed
to crew
TABLE 4.2
GROUNDTEST INSTRUMENTATION (Continued)
MEASUREMENT
NAME
MEASUREMENT
NUWGER*
RANGE
VALUE
SCO362U
Fan Motor Oper
Tank 1 H2
SCO363U
Fan Motor Oper
Tank 2 H2
SC2075X
FC H2 Inline
Htr ON
ON
Event
SC2076X
FC H2 Inline
Htr OFF
OFF
Event
SC2116V
FC 1 DC Volts
Out
25-40 volts
0.059
SC2117U
FC 2 DC Volts
Out
25-40 volts
0.059
sc2iiau
FC 3 DC Volts
Out
25-40 volts
0.059
SC213OX
FC 1 H2 Purge Valve
Oper
Close - Open
Event
SC2131X
FC 2 H2 Purge Valve
Oper
Close - Open
Event
SC2132X
FC 3 H2 Purge Valve
Oper
Close - Open
Event
* See note,
page A-160
TABLE 4.2
GROUNDTEST INSTRUMENTATION (Continued)
MEASUREMENT
NUMBER*
-
MEASUREMENT
NAME
BIT
VALUE
RANGE
I
SC2133X
FC 1 O2 Purge Valve
Close - Open
Event
sc2134x
FC 2 O2 Purge Valve
Close - Open
Event
SC2135X
FC 3 02 Purge Valve
Close - Open
Event
SC2326X **
FC 1 0 /H Shutoff
Valve ape{ Hold
Off - Hold
Event
Off - Hold
Event
FC 3 0 /H Shutoff
Valve apeG Hold
Off - Hold
Event
GC5000V
FC 1 Htr Voltage
Zone 1
O-120 vrms
0.472
GC500lV
FC 1 Htr Voltage
Zone 2
GC5002V
FC 1 Htr Voltage
Zone 3
SC2327X **
sc232ax
**
* See note,
page A-160
** Data also
displayed
to crew
TABLE 4.2
GROUNDTEST INSTRUMENTATION (Continued)
I~IEASUREMENT
NUMBER*
MEASUREMENT
NAME
RANGE
BIT
VALUE
O-120 vrms
0.472
O-120 vrms
0.472
FC 1 Htr Current
Zone 1
O-5 arms
0.0197
GC5010C
FC 1 Htr Current
Zone 2
O-50 arms
0.197
GC50llC
FC 1 Htr Current
Zone 3
O-5 arms
0.0197
GC5003U
FC 2 Htr Voltage
Zone 1
GC5004V
FC 2 Htr Voltage
Zone 2
GC5005U
FC 2 Htr Voltage
Zone 3
GC5006V
FC 3 Htr Voltage
Zone 1
GC5007U
FC 3 Htr Voltage
Zone 2
GC5008V
FC 3 Htr Voltage
Zone 3
GC5009C
* See note,
page A-160
TABLE 4.2
GROUNDTEST INSTRU~IENTATION (Continued)
MEASUREMENT
NUMBER*
MEASUREMENT
NAME
RANGE
BIT
VALUE
GC5012C
FC 2 Htr Current
Zone 1
O-5 arms
0.0197
GC5013C
FC 2 Htr Current
Zone 2
O-50 arms
0.197
GC5014C
FC 2 Htr Current
Zone 3
O-5 arms
0.0197
GC5015C
FC 3 Htr Current
Zone 1
O-5 arms
0.0197
GC5016C
FC 3 Htr Current
Zone 2
O-50 arms
0.197
GC5017C
FC 3 Htr Current
Zone 3
O-5 arms
0.0197
GC5019E
FC 1 Htr Power
O-5000 watts
19.7
GC5020E
FC 2 Htr Power
GC502lE
FC 3 Htr Power
* See note,
page A-160
TABLE 4.1 NOTE
The measurement identification
used in Table 4.1 consists
of seven
two letters
followed
by four numbers and one letter
as
characters:
shown below.
DISCRETE NUMBER
FUNCTIOiJAL SYSTEM CODE
MODULE CODE
CLASSIFICATION
F"'"""'"'"'
Ji
blodule
SC 9099P
Code-
The first
letter
designates
C
Command Module
G
GSE Auxiliary
S
Service
Function
Subsystem
The second letter
originates:
Discrete
and Checkout
location
by module:
Equipment
Module
Code
denotes
Electrical
C
the measurement
the subsystem
within
which
the measurement
Power
Number
Characters
three through
within
each system.
Measurement
Classification
The seventh
or type:
character,
six
are discrete
a letter,
denotes
numbers listed
measurement
C
Current
R
Rate
E
Power
T
Temperature
P
Pressure
U
Voltage
Q
Quantity
X
Discrete
sequentially
classification
Event
A-160
I
^_
._
”
..,,,.,_.
l.l
..ll-_-l‘._,~,-”
..-A
---_I
_---”
--_...-
_ -----__--
_-,llll-_-_ll._
5.0
FUEL CELL/CRYOGENIC SUBSYSTEMMALFUKTION PROCEDURES
The procedures
describe
the proper order and nature of emergency
steps the crew must perform to determine
the source of a fuel cell
or cryogenic
storage system problem/malfunction.
A Caution and Warning
alarm and light
or abnormal instrumentation
indication
is evaluated
by
a malfunction
procedure
logic diagram.
The logic diagrams enable the
crew to determine
the source of the problem and corrective
actions,
if
required.
Fuel cell shutdown and bus short isolation
(not related
to
Caution and Warning) procedures
are also presented as part of the
malfunction
procedures.
The procedures
are primarily
used as a guide for
locate a problem and are presented for the flight
the crew actions.
The source
Procedures.
of the data was CSM 108 (Apollo
the flight
crew to
monitor as a guide
12) Flight
to
Malfunction
A-161
-- ..1_---"
----- -,. --_- -~...- _ . I_
^l-.l--_~~---.-"".--_. ~- -.- ~_.-.ll-l_~_
-.-_
YMPTOM
340 pia
220 1111
REMARKS
PROCEDURE
1
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
I
.
..I
.I
L-
.--
.
..-
-..----F
^..-“~--.l
PROCEDURE
..-ll ~_
--
-_.,.
-EMARKS
~
-.---
.~.
SYMPTOM
REMARKS
PROCEDURE
~-164
~.
,.- .., .----. .- ---."I_....----I. _.. ."-. .._-- I -.--".^.._._...
~"_",-~
_-..___
--,...----.---
PROCEDURE
~-165
I
REMARKS
A-166
PROCEDURE
~-167
REMARKS
PROCEDURE
REMARKS
SYMPTOM
PROCEDURE
m
3-
I
-
a-,
REMARKS
A-170
I"e.-"--I...~-_ .-_-_
j( _I__,_*_.
_-__"__l__.---,_.---
--
_._.
_...
-._.w- .._
__---_____,,-_.-__.-__ ...I -..,,.
E
:
7.0
The fuel
isometric
filtration
FUEL CELL/CRYOGENIC SUBSYSTEMHARDWAREDESCRIPTION
cell/cryogenic
hardware description
includes
the subsystem
drawings,
fluid
schematics,
component descriptions
and
provisions.
Isometric
drawings locate operational
hardware:
tubing runs, sizes
A schematic drawing of the
and part numbers: and system interfaces.
Environmental
Control
System describes
the water and oxygen system
interfaces.
Fuel cell and cryogenic
storage system schematics
aid understanding
These schematics
are also used to reference
of the system plumbing.
to specific
hardware component descriptions.
Filtration
clearances
data describe
and the filters
the component
rating,size,
protected,
location
its minimum
and type.
Hardware descriptions
are intended for rapid reference
to the specific
physical
hardware affected
as a result
of a malfunction.
Fuel cell/
cryogenic
subsystem interactions
with interfacing
components and subsystems are clarified
by this background information.
The sources of the data included
North American Rockwell Uperational
Checkout Procedures
(OCP's), Pratt and Whitney Aircraft
Fuel Cell
Electrical
Power Supply-PC3A-2
Support Manual, dated February 1, 1969,
Pratt and Whitney Apollo Fuel Cell Component Descriptions,
and Beech
Aircraft
Corporation
Project Apollo Cryogenic Gas Storage Subsystem
Flight
Support Manual, dated September 6, 1968. The descriptions
are
applicable
through CSM-115 including
identified
hardware changes for
shown were current
and correct
as
CSM 112-115 . The configurations
of December 1969.
A-172
.___...
.-,-_
._
.--.
*
.
..-_
-...
_._-..
..--f-
._”
l--ll.-ll.
“--
.___
-^--
-.
._-I.
.
.”
--....
-
.-..
._“.--.
*___“_.-^
.--.
..--.
7.1
SYSTEM HARDWAREISOMETRIC
DRAWINGSAND SCHENATIC
A-173
FUEL CELL/CRYOGENIC SUBSYSTEM LOCATION IN
SERVICE MODULE
/
\
\
/
\
\
C/M-S/M?MBILICAL
FUEL CELL SHELF
\
O2 SUBSYSTEM
SHELF MODULE\
H2 TANKS-
H2 SUBSYSTEM
SHELF MODULE \
I
BEAM NO. 3
VIEW LOOKING INBOARD SECTOR IV
A-174
FUEL CELL SHELF INTERFACE
FUEL PUMP
F/C SHELF
LEVEL
4
/N2
REG
I
-7Y
/
BEAM
NO. 3
BEAM
NO. 4
GLY FILL (LOOP NO. 1)
ME 273-0036-0002
N; FILL ME 273-0036-0001
ME 27%0075-0004
EPS 02 VENT
ME 273-0041-0001
INTERFACES
Fl - H2 IN
F2 - O2 IN
F3 - O2 VENT
F4 - H,O OUT
F5 - N; FILL
F6 - H2 VENT
F7 - WATER GLYCOL OUT
F8 - WATER GLYCOL IN
A-175
-
_..
.
. .
.~.
._
_
..__
,,
-1__-.-._
-...
-
..__..
-_..
,_“_..
OXYGEN SUBSYSTEM SHELF MODULE
c-\.
-.
BEAM NO. 3
VALVE MODULE
-OS-S
(ECS 2)
Y
Ll!!iiJ
I
y-O2
I
BEAM NO. 4
02 '
A-176
HYDROGENSUBSYSTEMSHELF MODULE
-
H, TANK NO. 1
BEAM NO. 3
BEAM NO.
HS-3 (F/C 3)
HS-2 (F/C 2)
HS-1 (F/C 1)
RELIEF
H2
!I
I 4,
FUEL CELL VALVE MODULE
Ii2 VALVE MODULE/
EPS WATER GLYCOL RADIATOR TEMPERATURESENSOR LOCATION
FR(
TO RAumj
SIGNAL CONDITIONERS
sc2092TSC2091 T-
-v
BEAM 4
SC20
an
F/C 1IRAU.
OUT1iET
\
F/C 2 KAU. INLET
SC2092T
F/C 3 KAU. INLET
VIEW LOOKING
INBOi
(FUEL CELL SHE1 c *DE’
WATERGLYCOL SERVICE MODULE LINES
(F/C3)
V37-458020-143
(TZ)
(3/S)'
(Tl)
(F/C3)
V37-458056-25
(Tl)
(F/Cl)
V37-458056-21
(Tl)
(F/C2)
V37-458056-23
iUOIATOR
RREA
RADIATOR AREA
FUEL CELL 12
/A
C No. 3
I
FUEL CELL 13
'4
EFFECTIVE
IIN S/C 098 THRU106
EFFECTIVE 01( 5/C 107 MO WRS
LINE SIZE III INCHS(OUlSIM
DIMTER)
NOTE: WL
[email protected]%.S - .OZO IMCIES MTERIM
30-L CRES
l WL
THICKMSS - .035 MIXES (MTERIN
N ALLOY)
A-179
.a,-+:
"37-G
CRYOGENIC HYDROGENSERVICE MODULE LINES
ME 273-O(
BEAr' T
F/C Shelf
[21v37-454049-23
q v37-454030-45
&Jv37-454031-29
(l/4)
(l/4)
(l/4)
[zlv37-454049-55
(l/4)
m ~37-454031-43
(l/4)
Ql V37-454049-63
(l/4)
m ~37-454030-47
(i/4)
QV37-454049-25
(l/4)
02 Shelf
V37-454208-107
v37-454208-113
Area
Area
(l/4
(318
m EFFECTIVE ON S/C 098 THRU 106
q
EFFECTIVE ON SjC 107 AND SUBS
( ) LINE SIZE IN INCHES (OUTSIDE DIAMETER)
NOTE: WALL THICKNESS = .020 INCHES
v37-454030-39
v37-454049-19
( l/4 Jm
(l/4)(21
JRYGGENIC OXYGEN SERVICE XODULE LIIiES
V37-454207-115
/c-3-v37-454207-H3
(l/4)
(l/4)
Beam 2
c OS5
Beam 3
---
-LECS 02 LINES
(CM-SM Umbilical)
F/C Shelf
Area-
1
v37-454031-25(1/4)(l)
TJv17-454049-53(1/4)(2)
(?)V37-454049-49
(l)V37-454031-21
(l)V37-454031-39
(2)v37-454049-61
(21v37-454049-51
(l)V37-454031-23
(l)V37-454031-33
(2)V37-454049-57
II
02 Shelf
H2 Shelf
I F/C 3
Area-
Area-
(1) EFFECTIVE ON S/C 098 THRU 106
2 EFFECTIVE ON S/C 107 AND SUBS
I ) LINE SIZE IN INCHES (OUTSIDE DIAMETER)
NOTE: WALLTHICKNESS = .020 INCHES
I
II11
k
V37-454207-111
(l/4)
-v37-454207-109
(l/4)
--V37-454207-107
(l/4)
454207-105 (l/4)
-v37-454207-103(1/4)
-OS-5
(ECS-2) '
%OS-4 (ECS-1)
%S-3
(F/C-3)
qos-2
(F/C-Z)
OS-1 (F/C-l)
'
FUEL CELL NITROGEN SERVICE MODULE LIiJES
B
F/C Shelf
Area --
I
El v37-454&9-33(1/4)7a v37-454030-57(1/4)
m
-P
\
--
v37-454030-59(1/4)
Ql v37-'!54049-35(1/4
N2 Supply
i
(F5)
q
al
a
I,
v37-454030-51(m)
v37-454030-49(1/4)[email protected]+2'7(1/4)-
EFFECTIVE ON S/C 098 THRU 106
EFFECTIVE ON S/C 107 AND SUBS
( ) LINE SIZE IN INCHES (OUTSIDE
DIAMETER)
NOTE: WALL THICKNESS = .020 INCHES
A-187
A-184
V37-454240-104
ME282-0047-0040
V37-454240-103
V37-454240-102
ME284-0119-0001
7-454240-101
ME284-0115-0001
1
O2 TANK 3
V37-454225-11
V37-454240-101
ME282-0046-0002
HYDROGEN/OXYGEN
TANK SHELF MODULE- SECTOR r (SIDE VIEW - BEAM 6 SIDE)
EFFECTIVE ON CStq 112-115
A-lP!iY
.-... ..,
~.
._.. .-.-
---_.
._.__.. ,^-.-.“-l..~
.^._ . .“. ,.__
--.--_
_.x._...__ - _... -_-_
.-
_--_lll-
HYDROGENRELIEF (HR)
r
/
SECTOR VI
’
/-
._
V37-454240-12.5
V37-454203-1194
TO ECS'-
V37-454240- 15
V37-454240-l
FROM O2 TANK 2
HYDROGEN/OXYGEN
SERVICE MODULE LINES
EFFECTIVE ON CSM 112-115
k
V37-454240-20
V37-454240-21
HYDROGENTANK 3
SUPPLY TO VALVE
SHELF MODULE
FROMH2 TANK 3 SECTOR I
FORCSM 112-115
V37-454240-24
HYDROGENTANK 3 SUPPLY
INTERFACE WITH TANKS 1 & 2 - HYDROGEN
SHELF VALVE MODULE- SECTOR IV
EFFECTIVE ON CSM 112-115
A-18cj
...
-
,;
.-.. ~
.-- .,_^
-_, . _~__
...l.--.-.._..
-.---.._..-.,
__-._-._
.-_I
7.2
FUEL CELL COMPONENTDESCRIPTIONS
A-190
--___
__-.
I_
.____,-____.
-I
..”
._.._
I_.”
.__.
_”
.--.
--..-.
.___._,.---1--.-1”
.,..
-
_...._
^
_
A-1()1
CRYOG
NIC GASSTORAGE
SYSTFN
7.3
COMPhENT DESCRIPTIONS
A-192
_.
.
^...
“1
.
.^.
-.~
.-__
_-.....
l*.-l-
.“_
.
.
.
-_--.
__.__
.
t
-_~--..^-“--
I_._.
.-.
.
.
”
-
.^
_.
.ll.“l.-_^-
.”
-.-.
.
-l-._-^--_~_l__“-
CRYOGENIC STORAGE SYSTEM SCHEMATIC
A
E
El
D 1
F2
0 R 1
1
g5
M
s
NOTE:
0
Denotes component
descriptions
in
numerical
order
(See Pages7-56 through
7-78)
HzTANK
3
TABLE 7.3.1
CRYOGENIC GAS STORAGE SYSTEM INTERFACES
CGSS INTERFACES
ELECTRICAL INTERFACE
Pigtails
Pigtails
Hermetically
sealed pin
receptacle
Hermetically
sealed pin
receptacle
Connections
l/4" O.D.
(0.015 wall)
3/8" O.D.
(0.022 wall)
Vent Connections
l/4" 0.0.
(0.015 wall)
3/4" O.D.
(0.028 wall)
Relief
3/16" O.D.
(0.022 wall)
3116" O.D.
(0.022 wall)
l/4" 0.b.
(0.015 wall)
l/4" O.D.
(0.022 wall)
Feed Connections
l/4" O.D.
(0.015 wall)
l/4" 0-D.
(0.022 wall)
F/C Supply
l/4" O.D.
(0.022 wall)
l/4" O.D.
(0,022 wall)
l/4" O.D.
(0,022 wall
l/4" O.D.
(0.022 wall)
l/4" O.D.
(0.022 wall
l/4" O.D.
(0.022 wall)
l/4" O.D.
(0.022 wall
l/4" O.D.
(0.022 wall)
Valve
Modules
Tank Connectors
TANK INTERFACE LINE SIZES
Fill
Connections
Feed Connections
CRYOGENIC VALVE MODULE
Relief
Connections
Valve
Outlet
FUEL CELL VALVE MODULE
Feed Connections
F/C Supply
(2)
Connections
(3)
m&/-?-j
OXYGEN AND HYDROGEN STORAGE TANK
Electrical
Fill
Vent
Connector
q
Fill
Vent
Connector
Connector
Feed
Connector
Connector
H2 STORAGE TANK
q
02 STORAGE TANK
Each storage
tank consists
of two concentric
spherical
shells.
The annular
space between
them is evacuated
and contains
the thermal
insulation
system,
pressure
vessel
support,
fluid
lines
and the electrical
conduit.
The inner
The
shell,
or pressure
vessel
is made from forged
and machined
hemispheres.
pressure
vessel
support
is built
up on the pressure
vessel
from subassemblies
and provides
features
which transmit
pressure
vessel
loads to the support
The fluid
lines
and the electrical
lead line
exit
the pressure
assembly.
vessel
at its
top,
traverse
the annular
space and exit
the outer
shell
as
02, top of tank coil
cover;
H2, girth
ring
equator.
follows:
Structural
respectively.
listed
in
.,...-
".._.
and physical
parameters
are listed
in Tables
7.3.2
and 7.3.3,
Tank volumes,
with expansion
and contraction
data,
are
Tube sizing
is listed
in Table
7.3.5.
Table
7.3.4.
._--
_- _.._--.._l_-..-.
.~._...-.--.-..--"-.l..l"-l Lo---.------
..,-..11..-
Line
TABLE 7.3.2
CRYOGENIC TANK STRUCTURAL LIMITS
Hydrogen
Oxygen
Material
Ultimate
Strength,
psi
Yield Strength,
psi
Young's Modulus, psi
Creep Stress,
psi
5 Al-2.5
Sn ELI Ti
105,000
95,000
17 x lo6
71,200
Poisson's
0.30
Inconel 718
180,000
145,0006
30 x 10
No creep at
145,000
0.29
1.5
1.33
1.33
1.5
1.33
N.A.
53,000
110,000
Safety
Ratio
Factors
-
Ultimate
Yield
Creep
Design
Stress
Level,
psi
Proof
Pressure,
psia
400 psia
1357 psia
Burst
Pressure,
psia
450 psia
1537 psia
A-196
.__-_- _ ._ ,_^.___ _^ .._II________x.
_. "_l_l___,
.__
_..-___
_. ---...-- -...._-..
__
_,__-_-~_. ^_..
.- ._1.'-'_------L"II__L-.._.I.^
.I .^.. ___-
TABLE 7.3.3
CRYOGENIC TANK PHYSICAL PARAMETERS
Parameter
Hydrogen
Pressure
Vessel
Material
Inside Diameter
Wall Thickness
Outside
Outer
Oxygen
5Al-2.5
28.24
- Inches
- Inches
Diameter
- Inches
Sn ELI Ti
Inconel
25.06
004
.044 c:ooo
.059
28.328
25.178
718
.004
faooO
Shell
5Al-2.5
Material
Inside
Diameter
Outside
- Inches
Diameter
Wall Thickness
- Inches
- Inches
Sn ELI Ti
Inconel
750
31.738
26.48
31.804
26.52
.033 + .002
.020 + .002
Support
Flange
Diameter
Flange
Thickness
Bolt
Circle
37.966
- Inches
.070 i
- Inches
Diameter
- Inches
Number of Bolts
28.228
.OlO
.080 c .OlO
32.216
27.50
8
12
Annulus
Annular
Space - Inches
Insulation
Vacuum Level
1.705"
Vapor-cooled
and passive radiation
shields.
5 x lo-7
(TORR) - MM Hg
.653"
Vapor-cooled
shield
with oreloaded
insulation.
5 x 10-7
24 Days
24 Days
10
90 psi r 2.
75 psi
Spec
75.0 lb.
93.5 lb.
Actual
80.0 lb.
90.8 lb.
Average
Burst
Pump Down Time
Disc
Burst
Weight
Pressure
f 7.5
(Empty)
(Maximum)
Electrical/Instrumentation
Beech/NAA Interface
Pigtails
& Hermetically
sealed pin receptacle
A-197
Hermetically
sealed
pin receptacle
TABLE 7.3.4
CRYOGENIC TANK VOLUMES
(With Expansion
and Contraction
Data1
INTERNAL VOLUME
.02 ft3 for components
(Less
H2 Tank
02 Tank
Ambient
Pressure
Max.
tol.
4.7528
ft3
6.8314
ft3
Ambient
Temperature
Max.
tel.
4.7471
ft3
6.8045
ft3
Max.
Min.
tol.
tol.
4.7213
4.7156
ft3
ft.3
6.777 ft3
6.7698 ftd
Max.
tol.
4.7532
ft3
N/A
Min.
tel.
4.7497
ft3
N/A
Max.
tel.
4.7508
ft3
Min.
tol.
4.74705
Max.
tol.
Min.
tol.
N/A
N/A
Full Ambient
-297OF 02
-423OF
Pressure
H2
Full
935 psia
(02)
-294'F
Full
865 psia
(02)
-294'F
Full
260 psia
-418'F
Full
225 psia
(H2)
Full
935 psia
ft3
6.80048
ft3
--
tol.
WA
6.8014
ft3
tol.
N/A
6.7963
ft3
Max.
tol.
4.7848
ft3
N/A
Min.
tol.
4.7812
ft3
N/A
Max.
tol.
N/A
6.8597
ft3
Min.
tol.
N/A
6.8550
ft3
6.7988
Ft3
6.8573
Ft3
(02)
(H2)
-ZOOoF
Volumes
6.805
Max.
+ 80°F
Full
260 psia
N/A
Min.
(Hz)
-418'F
N/A
ft3
Used for
Tank Calculations
4.74892
Average Size
Cold
835 psia O2
225 psia H,
Ft3
L
4.7830
Average Size
Warm
935 psia 02
260 psia H2
Ft3
A-198
_I-~
.-
.._-
-.....
..“_
T..-I-
..----
--.-
_
,_
______..
_.
._
. . .
--I
_-.
”
. .._.--
.I
CRYOGENIC TANK TUBE SIZING
TABLE 7.3.5
Hydrogen
Oxygen
Vent Tube
l/4 O.D. x .015
wall 304 L SST
l/2 O.D.
(Inside
3/4 O.D.
(Outside
Inconel
Fill
l/4 0.0. x -015 wall
304L SST
3/8 O.D. x .022 wall
Inconel 750 AMS 5582
Common with
line
l/4 O.D. x .015 wall
Inconel 750 AMS 5582
Tube
Feed Tube*
Electrical
Tube
Vapor Cooled*
Shield Tube
vent
l/2 O.D. x .015 wall
304L SST
l/2 O.D. x .015 wall
Inconel 750 AMS 5582
l/4 O.D. x .015 wall
304L SST
3/16 O.D. x .015 wall
Inconel 750 AMS 5582
l/4 O.D. x .015 wall
Inconel 750 AMS 5582
Pressure Vessel to Vapor*
Cooled Shield Tube
* Three
tubes
joined
x .015 wall
coil cover)
x .028 wall
coil cover)
750 AMS 5582
to provide
a single
feed line
for
the oxygen
tank only.
A-199
.
.._.
--
.
-.
--.
_
,.
-----...
.-..-.
.s
_-
_,
~--_---
1_1
i-l-
-
._..___._--
___--.--
--11
_----
--*-----
-.
q
PRESSURIZATION AND DESTRATIFICATION UNIT
Each of the storage tanks con tains a forced
cation unit.
pressurization
and destratifi
Each unit consists
of the fol lowing:
convection
a.
A 2.0 inch diameter s upport tube
3/4 the tank diameter in length.
b.
Two heaters.
c.
Two fan motors.
d.
Two thermostats.
Eliminated
for H2 on CSM
113 and on and eliminated
for 02 on CSM
114 and on.
FAN & MOTOR
approximately
TUBE
The tube provides
a large surface area for efficient
heat transfer,
and is small enough to be installed
through the pressure vessel neck.
The heaters are
placed along the tube's outer surface
.* and brazedI . in
A motor-fan
is mounted at the upper ana lower
place.
which draw fluid
through the inlet
ports located along
across the heat transfer
surface and expel it near the
vessel.
Block II tanks utilize
separate sets of lead wires for
and for each motor fan through the electrical
connector
FAN & MOTOR
ENCASED
INTERNALLY
enas of the tube,
the tube, force it
top and bottom of the
each heater
interface.
element
FAN MOTORS
The motors are three phase, four wire, 200 volts A.C. line to line,
400
The
cycles miniature
induction
type with a centrifugal
flow impeller.
minimum impeller
speed of the oxygen unit in fluid
is 1800 rpm with a
torque of 0.90 in. oz., and the hydrogen unit is 3800 rpm with a torque
Two fans and motors are used in each vessel.
of 0.45 in. oz..
Stator Stack ,
Yoke Ring
Bushing
Field
/--Winding
Bearings -4
Impel ler
\
MOTOR FAN
A-200
Rotor
q
PRESSURIZATION
S’I’ATOR
STAMR
TEZl’H
AND DESTRATIFICATION
UNIT(COIITINUED)
SLOT
-\
SPIDER
STATOR
RFTAIRER
i
I-
R?XESiED
STRAIN
SWI’
RELIEF
A-201
. . _ .,
..).-....---. _-_-‘ -.- .."-.-____-_~
l___l___."_~-.-~_~..----
_-1__~
-_-__-_--
--11----
q
PRESSURIZATION AND DESTRATIFICATION UNIT
(CONTINUED)
HEATERS
The heaters are a nichrome resistance
type, each contained
in a
thin stainless
steel tube insulated
with powered magnesium oxide.
The heaters are designed for operation
at 28 volts DC during
in-flight
operation , or 65 volts DC for GSE operation
to provide
pressurization
within
the specified
time.
The heaters are
spiralled
and brazed along the outer surface of the tube.
The
heaters are wired in parallel
to provide heater redundancy at
half power.
The heaters have small resistance
variation
over a
temperature
range of + 80°F to -420°F.
THERMOSTATS
THRUST PIN
WAVE WASHER
.GLASS SEAL
BASE ASSEMBLY
INSULATOR
CAP
BI-METAL DISC
The thermostats
are a bimetal type unit developed for cryogenic
service.
They are in series with the heaters and mounted inside
the heater tube with a high conducting
mounting bracket arranged
so that the terminals
protrude
through the tube wall.
When the
heater tube reaches 80 *lOoF, the thermostats
open cutting
power to the heaters to prevent over heating of the pressure
vessel.
When the tube reaches -200°F in the hydrogen tank or -75°F in the
oxygen tank the thermostats
close allowing
power to
be supplied
to the heaters.
q
DENSITY SENSOR PROBE
The density
sensor consists
of two concentric
tubes which
serve as capacitor
plates,
with the operating
media acting
as the dielectric
between the two.
The density
of the fluid
is directly
proportional
to the dielectric
constant
and
therefore
probe capacitance.
The gage is capable of sensing
fluid
quantity
from empty to full
during fill
and flight
operation.
The accuracy of the probe is 1.5% of full
scale.
TEMPERATURESEN
DENSITY
PROBE
PRESSUREAND
OESTRATIFICATION UNIT fl
Hydrogen
Quantity
Gaging System
Range
Accuracy
Output Voltage
Output Impedance
Power
O-100% full
(.17-4.31
#/ft3)
~2.68 % full
range
O-5 V DC
500 ohms
2-l/2 watts 115 V
400 cps
q
O-100% full
O-5 V DC
500 ohms
2-l/2 watts
400 cps
115 V
TEMPEKATURESENSOR
The temperaturesensoris
a four-wire
platinum
resistance
sensing element
mounted on the density
sensor (see photograph of density
sensor probe).
It is a single
point sensor encased in a Inconel sheath which only dissipates
1.5 millivolts
of power per square inch to minimize self-heating
errors.
The resistance
of the probe is proportional
to the fluid
temperature and is accurate
to within
1.5%.
Hydrogen
Temperature Gaging System
Range
Accuracy
Output Voltage
Output Impedance
Power
-42O'F to -ZOOoF
t2.68 % full
range
O-5 VDC
5000 ohms
1.25 watts 115 V
400 cps
Oxygen
-32O'F to +80°F
t2.63 % full
range
o-5 v DC
5000 ohms
1.25 watts 115 V
400 cps
@
SIGNAL CONDITIONER
The temperature
and density
amplifiers
are separate modules, contained
The density
module functions
as an infinite
in the same electrical
box.
feedback balancing
bridge and utilizes
solid state circuitry.
The
temperature
module also uses solid state circuits
and amplifies
the
voltage
generated across the sensor which is linearly
proportional
to the
resistance
of the sensor.
The output in both cases is a O-5 volt DC
analog voltage which is fed into the NR interface.
The voltage
required
to run the signal
conditioner
is 115 V, 300 cycle single
phase, and
draws a total
of 3.75 watts of power.
The accuracy of the unit is 1.0%
of full
scale.
The modules
hermetically
are encased
sealed.
q
in Emerson-Cumings
epoxy potting
and the unit
is
ELECTRICAL CONNECTOR
The electrical
receptacle
is a hermetically
sealed device capable of
It contains
straight
pins
withstanding
system pressures
and temperature.
with solder cups attached.to
facilitate
the soldering
of lead wires from
units and the
the temperature
and density
probes, the destratification
The pins are sealed in a ceramic materi
al which has the same
heaters.
coefficient
of thermal expansion as the shell and pin material.
A-284
.-
.-._
“__
..”
-..“-._-_“~
.-..
~--.I
__.^
11-~
.
.~.
. ..----
T------
..,-I
-~
..-1_
1-1-1”^1.-1~
. . ..-
I
-..-
I
.”
--.-
1-1
..l--.l..
-I-..
.
. ..-
.*
q
VAC-10~ PUMP
DESCRIPTION
The vat-ion
pump is attached directly
to the vacuum annulus of the oxygen
tank which maintains
the insulation
space at reduced pressure required
for
adequate insulation.
Pumping action results
from bombarding the titanium
cathode with ionized
gas molecules which become chemically
bound to the
titanium.
The impacting
ions sputter
titanium
from the cathode.
The
sputtered
titanium
particles
also contribute
pumping by gettering
action.
The pump can be used as a vacuum readout device since the input
current
to the pump is directly
proportional
to pressure.
The unit is
powered by a DC-DC converter
capable of putting
out the required
amounts
of power.
CONSTRUCTION
Vat-ion
pumps have no moving parts.
The pumps consist
of two
titanium
plates spot welded to a vacuum tight
stainless
steel enclosure
with an anode structure
mounted between the plates connected to a coppergold brazed electrical
feedthrough.
A permanent magnet maintains
a
magnetic field
between the electrodes
causing the ions to follow
spiral
paths thus increasing
transit
time.
POWERSUPPLY ( CONVERTER)
The converter
is a solid state device capable of supplying
power to the vat-ion
The unit is energized
by a 28 V DC source
pump over a large range of pressure.
out 4.2 ma at 10 Volts
and is current
limited
to 350 ma. The unit is capable of puttinq
The unit employs a squ,are wave invert&r,
a
DC and lma at 4000 volts.
toroid
transformer
and a quadrupler
circuit
on the output.
Choke filters
are supp lied
on the 28 volt DC input to keep to acceptable
limits
the amount of
conducted interference
being fed back from the output.
The metal case is well
bonded to reduce to acceptable
limits
radiated
interference.
The circuits
are enc losed
in Emerson-Cumings stycast
2850 Ft.
A-205
q
VAC-ION
--
PUMP (CONTINUED)
PERFORMANCE
The pumping rate of the pump is constant
at 1 liter
is related
to pressure as shown by the graph below.
rnc>auKt
per second.
Pump current
VSlJJKKtNI
1 I/s VaclonPUMP
I
ma
Pump Current
LIFE SPAN
The practical
life
span of a vac- ion pump while
pressure ranges is as follows:
1
1
1
1
x
x
x
x
10e6
10-5
10-4
10-3
region
region
region
region
-
10,000
1,000
100
10
pumping in the various
hours
hours
hours
hours
A-20.6
_ .._
_I
.
‘ . "..-."-_"_
..--.-~. __.,-_ _. .._..
.---."._.(. . _---
q
FILTER
.FILTBR
The filter
is a multiple
disc type element rated at 175
The discs are stacked on a mandrel-like
cartridge.
The
to trap fibers
and particles
which could get downstream
hinder valve module and fuel cell operation.
The filter
inside the density
probe adapter and is welded onto the
line.
microns absolute.
filter
is used
of the tank and
is mounted
feed and vent
A-207
~~__r _"-_-^
I-.. ._._.__~..--. .--;"-_ -~ _-,___-
--
.-...~III_IIIIII-..-x--.---
_-_-
_-~I
q
SYSTEM (TANK) VALVE MODULE
h
TANK 2 HALF
TANK 1 HALF
CHECK
-,_m-----I
I____----
0R
0T
RELIEF VALUE
PRESSURE TRANSDUCER
OVERBOARDRELIEF
I
A
OVERBOARDRELIEF
0PS PRESSURE SWITCH
m CHECK VALVE
The system (tank) valve module for the hydrogen system and oxygen system
Each module contains
two relief
valves,
two
are functionally
indentical.
pressure transducers,
two pressure switches:
and one check valve.
These
module components are each separately
described
on the following
pages.
A-208
la
RELIEF VALVES
ATMOSPHERIC
SENSING
PORT-
NEGATIVE
RATE SPRING IL
ASSEMBLY
I
BELLOWS
POPPET
PRESSURIZED
VOLUME
POSITIVE RATE
SPRING ASSEMBLY
VENTA
L
TANK PRESSURE
The relief
valve, part of the system valve module, is differential
type
designed to be unaffected
by back pressure in the downstream plumbing.
The valve has temperature
compensation
and a self-aligning
valve seat.
The valve consists
of an ambient pressure sensing bellows preloaded with
a belleville
spring,
which operates a poppet valve.
Virtually
zero
pressure increase
between crack and full
flow is obtained
by cancelling
out the positive
spring rate of the pressure sensing element with a
negative-rate
belleville
spring
(see above right).
The large sensing
element and small valve produces large seat forces with a small crackto-reseat
pressure differential
assuring
low leakage at the reseat
pressure.
The Belleville
springs
are made of 17-4 PH and 17-7 PH
stainless
steels.
The bellows is a three-ply
device designed to prevent
fractures
due to resonant vibrations.
The relief
crack pressure
is 273 psig minimum for hydrogen tanks and
983 psig minimum for oxygen tanks.
The valve is atmospheric
sensing;
therefore,
relief
crack pressure in space is 273 psia minimum for hydrogen
and 983 psia minimum for oxygen.
Oxygen
Full
Flow Pressure
Reseat Pressure
Hydrogen
1010 psig
(max.)
285 psig
(max.)
965 psig
(min.)
268 psig
(min.)
A-209
.
i.
.I-
-...
.
.L-i-““_l.l”~-.._-l_-
.,.,.
_
-..--
.
--.
~__
-“,“111111
.“lll”----.,l---
_x__-_-
q
PRESSURESWITCH
TTANK
H
PRESSURE
rSENSING
DIAPHRAM
PIVOTED
TOGGLE LEVER
HORSESHOESPRING
REFE
PRES
INSULATOR '
A
.ELECTRICAL
CONTACT ARM
The pressure switch,
part of the system valve module, is a double pole,
single
throw absolute
device.
A positive
reference
pressure
(less than
atmospheric)
is used to trim the mechanical trip mechanism to obtain the
required
absolute
switch actuation
settings.
The reference
pressure is
typically
between 4 to 10 psia.
A circular
convoluted
diaphragm senses
tank pressure and actuates
a toggle mechanism which provides
switching
to drive motor switch (Cryogenic
Electrical
Control
Box Assembly).
The
motor driven switch controls
power to both the tank heaters and
destratification
motors.
The pressure switch body is 302 stainless
steel
the diaphragm is 17-7 stainless
steel.
This unit is capable
of carrying
the current
required
by the motor driven switch without
any
degradation.
The convoluted
diaphragm actuates
the switch mechanism in
a positive
fast manner which eliminates
bounce and the resultant
voltage
transients.
and
11
CRYOGENIC PRESSURETRANSDUCER
TANK
The pressure transducer,
part of the system valve module, is an absolute
'The transducer
consists
of a silicon
pickup
(vacuum reference)
device.
comprised of four sensors mounted on a damped edge diaphragm and an
integral
signal
conditioner.
The unit senses tank pressure through the
discharge
line from the tank.
The signal conditioner
output is a O-5
VDC analog output which is linearly
proportioned
to tank pressure.
Hydrogen
Range
Accuracy'
Output Voltage
Output Impedance
Power
Voltage
0 to 350 psia
+ 2.68 % full
O-5 V DC
500 ohms
1.5 watts
23 V DC
Oxygen
range
50 to 1050 psia
* 2.63 %full
range
O-5 V DC
500 ohms
1.5 watts
23 V DC
A-211
,.
, ._..~_..-
.._ .-...._.- -. I ,_..i _-...
-.--.-L
,_.,__-.^ ..I_.. --,.. _...-*....^..
- _ .. ..- _..I,._L.‘ _
_~--l_l-__-__--..,
-_.~~.._-l_-,.l
~I-.
•l
CHECK VALVE (SYSTEM MODULE)
From Tank 2
t
G?7
Spring
Seat Assv -
Seal
The check valve,
part of the system valve module, is designed to open at
The single
poppet IS
a differential
pressure of approximately
1 psia.
during flow in
spring loaded and has a large area to prevent chattering
This large area also helps in obtaining
a positive
the normal direction.
seal if pressurized
in the reverse direction.
A- 2.12
El
FUEL CELL VALVE MODULE
VALVE MODULE ENVELOPE
TO FUEL CELLS
-FROM
TANK
FLOWSCHEMATIC
The fuel cell valve module consists
shutoff
valves contained
in a cast
modules are functionally
identical.
are described
on succeeding pages.
of two check valves and three solenoid
The separate hydrogen and oxygen
body.
Individual
valve module components
A-213
.
__
.
_“-.*
.--
*.-._--...
,..
.._.
____-“-^
‘_--.”
-.-..
---
-.....
.‘_“..,...I
-...-
--
q
SOLENOID VALVES
POSITION SWITC
SOLENOID NO. 1
OUTLET
SIMPLIFIED
The solenoid
valves,
part of the fuel cell valve module, employ a poppetThis poppet is actuated
by a magnetic armature which is
seat arrangement.
The upper solenoid
is used to open the
suspended on a Belleville
spring.
The snap-over-center
belleville
spring both
valve;
the lower to close it.
guides the armatures and latches
the valve open or closed.
A switch to
The valve opens against
indicate
valve closed position
is incorporated.
pressure and pressure helps seal the valve against leakage in the normal flow
The valve body is 321 stainless
steel.
The maximum in-rush
direction.
current
is 10 amps with steady state current
circuit
has diode
noise
suppression.
A- 214
at 2 amps. The solenoid
coil
q
f
CHECK VALVE (FUEL CELL MODULE)
FROM SYSTEM
VALVE MODULE
MAIN S
AUXI
SEAT
Seated
auxiliary
- both main and
seats are closed.
The check valve,
part of
differential
pressure of
main seat and 'auxiliary
large seat area provides
the reverse direction.
Cracked
- at low flows
the auxiliary
seat is
barely
open and catches
contaminant
particles,
the main
seat is wide
open and protected
from
contaminants.
Full flow - both main
and secondary
seats are
wide open; the high flow
velocities
carry
particles
through
the valve
without
fouling
the seat.
the fuel cell module, is designed to open at a
approximately
1 psia.
The valve consists
of a
seat operating
as shown pictorially
above.
A
a positive
low leakage seal if pressurized
in
8 ti2-02
Ezl
INLINE FILTER
.-+qfq!p-&
-
FLOW DIRECTION
The hydrogen and oxygen reactant
filter
consists
of a multiple
of chemically
The discs are stacked on a mandrel-like
cartridge.
The filter
etched discs.
is used to trap contamination
which could get downstream of the reactant
tank
The filter
is rated at 5~ nominal and 12~ absolute
with a
valve modules.
The filter
design does not allow it to
dirt
holding
capacity
of .25 grams.
generate system contamination
and provides
closer adherence to specified
filter
rating.
A-216
--_ ".,_.l._-~-.l----l--_--.l-,-~-_
.-"---. ,._--- -,.,
--__.-_ -..--ll.l".lll .._
_,._-4__
-,"-..--.-.,----.--w--e.-.--
I_._.
q
FILL AND VENT DISCONNECTS - AIRBORNE
Each vent and fill
disconnect
utilizes
a spring loaded poppet and a pressure
cap that can be locked into place.
The ground unit is connected by aligning
grooves on the ground sleeve with keys on the airborne
body, pushing until
a
stop is reached (about 40 lbs. force is required),
and turning
the ground
sleeve until
engagement is complete.
The spring loaded poppets can be self
opening on installation
of mating ground disconnects,
or can be opened subsequent to installation
of the ground disconnect,
depending on the type of
ground unit that is used,
The poppet is self closing
on removal of the
ground unit regardless
of the type used.
A- 21’7
. ..-
.
,
_.
.._
^.”
,.“__-___
.
__*_*_.
_l^--~.-l_---.“~---~-~
------
-il_-.
_“-ll-“l_-~-
--.-
__L.-_-_
-___
_-
._-_-_
7.4
FUEL CELL/CRYOGENIC
SYSTEM FILTRATION
A-218
_
I
.--.
_,..- _ .
". ^..^--..---".-_- I .--_--~_-. ,__-,..____-..
,_____..
.
.-.^_( ,I__..-I- _--.~--__
TABLE 7.4.1
FUEL CELLS/CRYOGENICS - FILTRATION
MINIMUM
CLEARANCE
CRITICAL
COMPONENT
acondary
Bypass Valve
ater/Glycol
Pump
0 to 0.006 in.
tapered pintle
on travel
2 Pump
rimary
Bi-Metalic
Flapper
Stroke = 0.040 in.
Clearance (min) at full
regeneration
= 0.013 in.
Bypass Valve
No filter
annulus
depending
Hole Size = 0.030 in.dia
Stem Clearance = 0.013 in.
Max. Stroke = 0.048 in.
ater Separator
heck Valve
75~ nom.
100~ absolute2
Area = 6.6 in
Internal
to
Pump Inlet
I
Non Bypassing
Type
40~ absolute
Area2= 0.076
in.
Internal
to Water
Separator Pump
Made from
Sinter Cd
Powder
5~ nom.
12~ absolute
Holding capacity =
.25 grams
1
5p nom.
12~ absolute
Holding capacity =
.25 grams
Between H -0 Valve
Module & fi -a2 Fuel
Cell Moduli
1
Chem Milled
Stacked Disc
Filter
Element
Between H -0 Valve
Module & 6 -a2 Fuel
Cell Modulz
Chem Milled
Stacked Disc
Filter
Element
D
1
D
2 Regulator
OTHER
CHARACTERISTICS
FILTER PROTECTION
LOCATION
RATING - SIZE
Valve seat clearance
open = 1Op nom.
0.008 in. Min. radial
slid25~ absolute2
Area 0.076 in
ing clearance
exposed to
gas = 0.006 inches
D
1
D
Internal
Regulator
to
Inlet
Made from
Sinter Powder
I9
D
D2
Filtration
Valve
for
also
open seat
Apollo
provided
clearance
8 regulator.
by filter
at the H2 Regulator
is based on regulator
flow
Inlet
conditions
(see H2 Regulator)
at 2200 watts
plus
purge
TABLE 7.4.1
FUEL CELLS/CRYOGENICS - FILTRATION (Continued)
MINIMUM
CLEARANCE
CRITICAL
COMPONENT
H2 Purge Valve
1. H2 Purge Valve
Orifice
(Valve
exit)
611 nom.
18~ absolute2
Area 0.35 in.
Internal
to
Valve Inlet
CylindricalShaped Screen,
Bypassing Type
0.0305
Protective
screen (perforated
cap)
hole size =
0.008 in.
Internal
and
Upstream of Orifice
@ Valve Exit
Lee Jet Size
0.0305 in.
orifice,
750
LOHM
in.
dia.
Valve Seat clearance
open
= 0.005 in. Min. radial
sliding
clearance
exposed
to gas = 0.0035 in. b
02 Purge Valve
Ball Travel from Seat =
0.020 in. to 0.025 in.
Min. diametr c clearance
0.005 in.
open
I% Valve
for Apollo
OTHER
CHARACTERISTIC!
FILTERt PROTECTION
LOCATION
RATING - SIZE
Ball travel
from seat =
0.020 in. to 0.025 in.
Min diametric
clearance
=
0.005 in.
02 Regulator
1. O2 Purge Valve
Orifice
(Valve
exit)
1
0.0120
seat clearance
8 regulator.
in.
Internal
10~ nom.
?egulator
25~ absolute
Area = 0.076 in 21
to
Inlet
lade from
jinter
Powder
Internal
to
6~ nom.
\Jalve Inlet
18~ absolute
= Area = 0.35 in2
d a.
is based on regulator
Protective
screen (perforated
cap)
hole size =
0.008 in.
flow
conditions
IylindricalShaped Screen,
3ypassing Type
Internal
and
llpstream of Orifice
P Valve Exit
at 2200 watts
plus
Lee Jet Size =
3.0120 in.
orifice,
4500
LOHM
purge
TABLE 7.4.1
FUEL CELLS/CRYOGENICS - FILTRATION (Continued)
MINIMUM
CLEARANCE
CRITICAL
COMPONENT
2 Vent Valve
1. N Vent Valve
Okfice
Ball seal .020 to .025 inches
from seat. Valve pintle
travel
from sealing
seat is .OlO-.012
Maximum diametric
in.
clearance
.006 inches.
6~ nom.
Internal
to
Valve Inlet
18~ absolute
Area = 0.35 in?
0.0186
Protective
screen (perforated
cap)
hole size =
0.008 in.
in.
dia.
2. N Vent Valve
V&t Port Plug
I2 Fill
Valve
1. Inlet
Port
Port
I2 Regulator
I3
CylindricalShaped Screen,
Bypassing Type
Internal
and
Upstream of Orifice
0 Valve Exit
Lee Jet Size =
0.0186 in.
orifice,
2000
LOHM
6~ nom.
Internal
to
Valve Exit
18~ absolute
Area = 0.15 in?
Cylindrical
Screen
6~ nom.
Internal
to
18~ absolute
Valve Inlet
Area = 0.034 in?
(Min)
Disc-shaped
Screen
6~ nom.
Internal
to
18~ absolute
Valve at InterArea = 0.074 in?stage and Exit
(Min)
Port
Cylindrical
Screen
Ball seal .020 to .025 inches
from seat. Valve pintle
travel
from sealing
seat is .OlO-.012
in.
Minimum diametric
clearance
.005 inches.
2. Interstage
3. Exit
OTHER
CHARACTERISTICS
FILTER PROTECTIO!i
LOCATION
RATING - SIZE
Valve Seat Clearance Open,
0.0015 in. Min. radial
sliding
clearance
exposed
to gas = 0.0035 in.
I9
Valve open seat clearance
for Apollo 8 regulator.
is based on regulator
10~ nom.
25~ absolute
Area 0.076 in.'
flow
conditions
Internal
Regulator
to
Inlet
at 2200 watts
Made from
Sinter Powder
plus
purge
TABLE 7.4.1
FUEL CELLS/CRYOGENICS - FILTRATION (Continued)
MINIMUM
CLEARANCE
CRITICAL
COMPONENT
1 Regulator
Overboard
J&t Port Plug
rletering
0.0215
Orifice
Radiator Bypass Valve
Main Valve
Bypass Valve
Orifice/Stroke
Main Valve
Bypass Valve
in.
dia.
0.0015 to 0.003 in.
0.002 to 0.004 in.
0.156
0.420
in/O.060
in/O.018
in.
in.
FILTER PROTECTION
LOCATION
RATING - SIZE
200 MESH screen
0.001%0.0026in
dia.wire
Internal
Regulator
to
Exit
Protective
screen (perforated
cap)
hole size =
0.008 in.
Between N2
Regulator
and 02 &
H2 Regulators
None at Valve
Provided by
Pump Inlet
Filter
Filtrati'on
at Coolant
Inlet
OTHER
CHARACTERISTICS
Screen Twill
Plain
or
None
Filter
Pump
Valve Module, H2
(Consists
of 2 relief
lalves,
2 press.
iwitches,
2 press.
transducers,
and one
:heck valve)
175~ absolute
Internal
to H2
Area = 0.97 in? Cryogenic Tanks
Outlet
Chem Milled
Stacked Disc
Filter
Element
Jalve Module, O2
(Consists
of 2 relief
valves, 2 press.
switches,
2 press.
transducers,
and one
check valve)
175~ absolute
Internal
to 02
Area = 0.97 in? Cryogenic Tanks
Outlet
Chem Milled
Stacked Disc
Filter
Element
<
I
TABLE 7.4.1
FUEL CELLS/CRYOGENICS - FILTRATION (Continued)
CRITICAL
COMPONENT
12-0 Fuel Cell Valve
lodu?e (Consist of 2
:heck valves and 3
solenoid valves each)
MINIMUM
CLEARANCE
-T
FILl
RATING - SIZE
5~ nom.
12~ absolute
iolding
capacity =
.25 grams
PROTECTION
LOCATION
3etween H -0 Valve
'lodule ani H2-0
-uel Cell Mo&l$
OTHER
CHARACTERISTIC
Chem Milled
Stacked Disc
Filter
Element
Was this manual useful for you? yes no
Thank you for your participation!

* Your assessment is very important for improving the work of artificial intelligence, which forms the content of this project

Download PDF

advertisement