APPENDIX A - VEHICLE AND EQUIPMENT DESCRIPTION
This section contains a discussion of changes to the spacecraft, the
extravehicular systems, and the scientific equipment since Apollo 14. In
addition, equipment used on Apollo 15 for the first time is described.
The Apollo 15 command and service module (CSM-112) was of the block II
configuration, but was modified to carry out a greater range of lunar
orbital science activities than had been programmed for any previous
mission. The lunar module (LM-10) was modified to allow an increase in lunar
surface stay time and accommodate a larger scientific payload. The launch
escape system and the spacecraft/launch vehicle adapter were unchanged.
The Saturn V launch vehicle used for this mission was AS-510. The significant
configuration changes for the launch vehicle are given in reference 1.
A.1 COMMAND AND SERVICE MODULES
A.1.1 Structure and Thermal Systems
A scientific instrument module was installed in sector I of the
The module containing instruments for the
acquisition of scientific data during lunar orbit was attached with 1/4-
inch bolts to radial beams 1 and 6, to the new cryogenic tank panel, and to
the aft bulkhead of the service module. The sides of the scientific
instrument module were constructed of aluminum stiffened sheet, and the
shelves that supported the instruments were made of bonded aluminum
sandwich. A door covered the module until about 4 1/2 hours prior to lunar
orbit insertion when it was pyrotechnically cut free and jettisoned in a
direction normal to the X-axis of the spacecraft
covers and thermal blankets provided thermal control for individual
instruments within the module. For additional thermal control, the inside
surfaces of the module were coated with a material having an absorptivity-
to-emissivity ratio of 0-3/0.85; the surfaces facing the radial beams, and the
radial beams themselves, were coated with a material having an absorptivity-
to-emissivity ratio of 0.05/0.4. The instruments are discussed; in section
Because of the requirement to retrieve film cassettes from the
scientific instrument module during transearth coast, extravehicular
activity handrails and handholds were installed along the sides of the
module and inside the scientific instrument module. A foot restraint was
also attached to the module structure
A.1.2 Cryogenic Storage
A third hydrogen tank was installed in sector I of the service
module, as planned for all J-type missions. The isolation valve
between oxygen tank 2 and 3 was moved from sector IV to the
forward bulkhead to
decrease its vulnerability in the event of a catastrophic tank failure.
All single-seat check valves in the hydrogen and oxygen lines
were replaced with double-seat valves having greater reliability.
Thermal switches formerly used in the hydrogen tank heater circuits inside the
tanks were removed.
A scientific data system was integrated with the existing telemetry
to provide the capability for processing, storing, and
transmitting data from the scientific instrument module. The data
processor, located in the scientific instrument module, necessitated
changes to the data storage equipment and the introduction of a data
modulator and a tape recorder data conditioner. The data storage equipment
was modified to have twice the recording time of the previous equipment,
and was redesignated the data recorder reproducer. The tape recorder data
conditioner was added to minimize flutter-induced jitter of recorded pulse-
A.1.4 Displays and Controls
Switch S30 was deleted from panel 2 and its function was incorporated
into switch S29 so that both cabin fans operated simultaneously. Toggle
switch S137 was added to panel 2 for hydrogen tank 3 fan motor control. The
pressure and quantity outputs of hydrogen tank 3 were connected to meter
displays through switches S138 and S139 on panel 2. Panels 181 and 230 were
added to provide controls for the experiment equipment in the scientific
instrument module. Experiment cover controls were added to Panel 278. Panel
was added to provide umbilical connections for
extravehicular activity. Panel 604
was added to provide an audio
warning signal to the extravehicular crewman in the event of low suit
pressure or low oxygen flow.
The diameter of the fuel inlet orifice in the service propulsion
system was decreased to improve the propellant mixture ratio.
A.1.6 Environmental Control System
Several oxygen components were added to accommodate the scheduled
extravehicular activity for retrieval of data from the scientific
instrument module. The command module components consisted of a larger
restrictor and filter for the higher flow rate, check valves to prevent
backflow, connectors for the attachment of the umbilical, and a pressure
A.1.7 Crew Provisions and Extravehicular System
The Command Module Pilot's space suit was basically the same as the
Apollo 14 lunar surface suits except that the water connector and lunar
module attach points had been removed. An umbilical assembly (fig. A-6) was
furnished to serve as a tether and provide oxygen, communications, and
electrocardiogram and respiration rate measurements for the extravehicular
crewman. An adapter plate mounted on the chest of the suit allowed
attachment of an oxygen purge system (transferred from the lunar module).
The purge valve was also brought from the lunar module to be used with the
oxygen purge system. A pressure control valve was provided to maintain suit
pressure at 3.5 to 4.0 psia at a flow rate of 10 to 12 lb/hr during the
extravehicular activity. A suit control unit
was connected to
the suit end of the umbilical to maintain the desired oxygen flow rate and
activate the suit pressure alarm if an anomalous condition had been sensed.
An 8-foot tether was furnished for use by the intravehicular crewman
stationed at the hatch (fig. A-5). The tether prevented forces from
being applied to his oxygen umbilical. In
addition, a thermal cover was furnished to protect his communications
An extravehicular activity monitor system was furnished to allow
television and 16-mm camera coverage of the extra-vehicular crewman's
activities. The components the system consisted of a sleeve mount attached
to the side hatch handle and a 34-inch pole assembly to mount the cameras.
A.2 LUNAR MODULE
A.2.1 Structure and Thermal Systems
A number of structural changes were made to the lunar module in order
to provide greater consumables storage capacity, permit stowage of a lunar
roving vehicle, and allow a heavier load of scientific equipment to be
carried. The most significant structural changes were as follows:
a. The descent stage propellant tanks and the openings for the
tanks were enlarged.
Heaters, additional insulation and shielding were incorporated in
quadrants I, III, and IV of the descent stage to protect equipment stowed in
those areas. Insulation in the docking tunnel was increased, and shielding
was added to reduce the heat leak to the cabin through the docking tunnel.
The fire-in-the-hole shield as well as the base heat shield were modified to
accommodate changes in the descent propulsion system (par. A.2.4).
b. Two tanks and supporting structure were added in descent stage
quadrant IV for storage of water and gaseous oxygen.
c. The structure in descent stage quadrants I and III was modified to
accommodate the lunar roving vehicle and its equipment pallet,
d. The descent stage beam panels, tank supports, lower diagonals,
beam, capstrips, and the ladder were strengthened structurally.
e. Descent batteries 1 and 2 (previously located in quadrant IV)
and descent batteries 3 and 4 (previously located in quadrant I) were
moved to the minus Z outrigger.
f. The size of the modular equipment stowage assembly was increased.
The ascent stage reaction control system tanks were insulated, and the
coating on the tank bay thermal shields was changed to a material with a
lower absorptivity-to-emissivity ratio to compensate for the extended lunar
stay time and higher sun angles.
A.2.2 Electrical Power
In addition to the four descent batteries (par. A.2.1), a fifth
battery (called the lunar battery) was provided to increase lunar stay time
capability. The capacity of each battery was 415 ampere-hours compared with
400 ampere-hours for previous missions. Other differences in the descent
batteries were as follows:
a. The battery relief valve, cell manifold relief valve and
pressurizing port adapters were changed from nylon plastic to ABS
A battery relay control assembly was added to route battery status
information to the proper channels because of the electrical control
assembly sections shared by batteries 2, 3, and the lunar battery, and an
interlock was added so that the lunar battery could not be switched to
both buses at the same time.
b. The method for attaching the cell manifold to the manifold
relief valve adapter was changed to prevent leakage.
A-2.3 Instrumentation and Displays
Water sensors were changed from quantity measuring devices to pressure
transducers for greater reliability. Descent fuel and oxidizer temperature
sensors were changed from immersion to container-surface measurements
because the measurements would provide more useful data. Temperature sensors
were added in the modular equipment stowage assembly to provide flight
statistical data. Instrumentation was added, and controls and displays were
changed on panel 14 because of the addition of the lunar battery.
The descent propellant system was modified to increase the tank capacity
1200 pounds, and the engine performance and operating life were increased.
These changes involved: (1) increasing the length of the tanks, (2) changing
material in the thrust chamber from an ablative silicon to an ablative
quartz, (3) replacing the exit cone with a lightweight cone, and (4)
increasing the nozzle extension 10 inches. Routing of pressurization lines
was modified to accommodate the larger propellant tanks. Modifications to
decrease the amount of unusable propellant consisted of deleting propellant
balance lines between like tanks and adding trim orifices to the tank
discharge lines (one orifice is fixed and the other is adjustable).
The oxidizer lunar dump valve installation was modified to be
identical to the Apollo 14 fuel lunar dump valve configuration. Thus, both
valves were installed to reverse flow direction through them and an
orifice was added upstream of each valve. This change was made to insure
that the valve would remain open with either liquid or gas flow.
In the reaction control system, a weight reduction of approximately
25 pounds resulted from the removal of the isolation valves from all
A.2.5 Environmental Control System
Extended stay time on the lunar surface required an increase in the
supply of lithium hydroxide cartridges. The oxygen and water supply was
increased for the same reason by adding a storage tank in the descent
stage for each system. Check valves were added at the outlets of the
original and new tanks, and servicing quick disconnects and pressure
transducers were added in association with the new tanks.
A new high pressure (approximately 1400 psia) portable life Support
system recharge capability was incorporated in conjunction with the added
oxygen tank. The recharge assembly includes regulators, overboard relief
valves, an interstage disconnect, a shutoff valve, and a quick disconnect to
mate with the portable life support system recharge hose. In addition, the
recharge hose was lengthened by 10 inches to permit recharging of the
portable life support system before it was doffed.
Instead of providing stowed urine bags and a portable life support
system condensate container as on Apollo 14, a 5-gallon tank was installed
in quadrant IV of the descent stage for both urine and portable life
support system condensate.
A.2.6 Crew Provisions and Cabin Stowage
Neck ring dust covers were provided to keep lunar dust out of the
pressure garment assemblies when not being worn. Tool carriers, attachable
to the portable life support system, were provided to facilitate carrying
of geological tools, sample bags and rock bags. An adapter was stowed to
permit the crewmen to connect their liquid cooling garments to the lunar
module water supply after removal of their pressure garment assemblies.
The ascent stage lower midsection and the lower left- and right-side
consoles were modified to carry additional lunar samples (each area could
carry a 40-pound bag). In order to carry the 70-mm, camera with 500-mm lens
and 70-mm film magazines, a special multipurpose container was installed in
the area behind the engine cover.
A-3 LUNAR SURFACE MOBILITY SYSTEMS
A-3.1 Extravehicular Mobility Unit
The pressure garment assembly was changed to improve mobility and
visibility, to permit easier donning and doffing, and to improve it
otherwise. The changes were as follows:
a. Neck and waist joints were added.
The portable life support system was modified to extend the lunar
surface stay time capability. There were four major changes:
b. The wrist rings were enlarged.
c. The shoulder area was modified.
d. The torso zipper was moved.
e. Gas connectors were repositioned.
f. A manual override relief valve was added.
g. The insuit drinking device was redesigned to hold 32 ounces of
water instead of 8 ounces.
a. An auxiliary water bottle was added.
A.3.2 Lunar Roving Vehicle
b. A larger battery was incorporated.
c. A higher pressure oxygen bottle was used.
d. Higher capacity lithium hydroxide cartridges were used.
The lunar roving vehicle
fig. A-7), used for the first time on Apollo
15, is a four-wheeled manually-controlled, electrically-powered vehicle that
carried the crew and their equipment over the lunar surface. The increased
mobility and ease of travel made possible by this vehicle permitted the crew
to travel much greater distances than on previous lunar landing missions. The
vehicle was designed to carry the two crewmen and a science payload at a
maximum velocity of about 16 kilometers per hour (8.6 mi/hr) on a smooth, level
surface, and at reduced velocities on slopes up to 25 degrees. It can be
operated from either crewman's position, as the control and display console
is located on the vehicle centerline. The deployed vehicle is approximately
10 feet long, 7 feet wide and 45 inches high. Its chassis is hinged such that
the forward and aft sections fold back over the center portion, and each of
the wheel suspension systems rotates so that the folded vehicle will fit in
quadrant I of the lunar module. The gross operational weight is approximately
1535 pounds of which 455 pounds is the weight of the vehicle itself. The
remainder is the weight of the crew, their equipment, communications
equipment, and the science payload.
The wheels have open-mesh tires with chevron tread covering 50 percent
of the surface contact area. The tire inner frame prevents excessive
deflection of the outer wire mesh frame under high impact load conditions.
Each wheel is provided with a separate traction drive consisting of a
harmonic-drive reduction unit, drive motor, and brake assembly. A decoupling
mechanism permits each wheel to be decoupled from the traction drive,
allowing any wheel to "free-wheel." The traction drives are hermetically
sealed to maintain a 7.5-psia internal pressure. An odometer on each
traction drive transmits pulses to the navigation signal processing unit
at the rate of nine pulses per wheel revolution. The harmonic drive
reduces the motor speed at the rate of 80:1 and allows continuous
application of torque to the wheels at all speeds without requiring
gear shifting. The drive motors are 1/4-horsepower direct-current,
series, brushtype motors which operate from a nominal input voltage
of 36 Vdc. Speed
control for the motors is furnished by pulse-width modulation from the
drive controller electronic package. The motors are instrumented for
thermal monitoring and the temperatures are displayed on the control and
is suspended from each wheel by a pair of parallel triangular arms
connected between the vehicle chassis and each traction drive. Loads
are transmitted to the chassis through each suspension arm to a
separate tension bar for each arm. Wheel vertical travel and
rate of travel are limited by a linear damper connected between the
chassis and each traction drive. The deflection of the suspension system and
tires combines to allow 14 inches of chassis ground clearance when the
lunar roving vehicle is fully loaded and 17 inches when unloaded.
Steering is accomplished by two electrically-driven rack and pinion
assemblies with each assembly steering a pair of wheels. Simultaneous use
of both front and rear wheel steering results in a minimum turning radius
of 122 inches. Steering is controlled by moving the hand controller left
or right from the nominal position. This operation energizes the separate
electric motors, and through a servo system, provides a steering angle
proportional to the position of the hand controller. The front and
rear steering assemblies are electrically and mechanically independent of
each other. In the event of a malfunction, steering linkage can be
disengaged, and the wheels centered and locked so that operations
can continue using the remaining active steering assembly.
Speed control is maintained by the hand controller. Forward movement
proportionately increases the forward speed. A neutral deadband
exists for about the first 1.5 degrees of forward motion. A constant
torque of about 6 inch-pounds is required to move the hand controller
beyond the limit of the deadband. To operate the vehicle in reverse,
the hand controller is pivoted rearward. However, before changing
forward or reverse directions, the vehicle must be brought to a full stop
before a commanded direction change can be made. Braking is initiated
in either forward or reverse by pivoting the hand controller rearward
about the brake pivot point. Each wheel is braked by conventional brake
shoes driven by the mechanical rotation of a cam in response to the hand
The vehicle is powered by two silver-zinc batteries, each having a
nominal voltage of 36 Vdc and a capacity of 120 ampere hours. During
lunar surface operations, both batteries are normally used simultaneously
on an approximate equal load basis. These batteries are located on the
forward chassis and are enclosed by a thermal blanket and dust covers.
The batteries are monitored for temperature, voltage, output current, and
remaining ampere hours on the control and display panel. Each battery is
protected from excessive internal pressures by a relief valve set to open
at 3.1 to 7 psi differential pressure. The circuitry was designed so that if
one battery fails, the entire electrical load can be switched to the
The control and display console is separated into two main functional
parts - navigation on the upper part and monitoring controls on the lower
part. Navigation displays include pitch, roll, speed, heading, total
as well as the range and bearing back to the lunar module. Heading is
obtained from a sun-aligned directional gyro, speed and distance
from wheel rotation counters, and range and bearing are computed from these
inputs. Alignment of the directional gyro is accomplished by relaying pitch,
roll and sun angle readings to earth where an initial heading angle is
calculated. The gyro is then adjusted by slewing with the torquing switch
until the heading indicator reads the same as the calculated value. The
displays utilize a radioluminescent material (promethium) that provides
visibility under lunar shadow conditions.
Thermal control devices are incorporated into the vehicle to maintain
temperature sensitive components within the necessary temperature limits.
They consist of special surface finishes, multilayer insulation, space
radiators, surface mirrors, thermal straps, and fusible mass heat sinks.
The basic concept of thermal control for forward chassis components is to
store energy during operation, and transfer energy to deep space while the
vehicle is parked between sorties. The space radiators are mounted on the
top of the signal processing unit, the drive control electronics, and on
batteries 1 and 2.
A-3.3 Extravehicular Communications
Because the lunar roving vehicle takes the crew beyond the range of
reliable radio communications with the lunar module using the portable life
support system communications equipment, radio communications equipment are
provided on the lunar roving vehicle that operate independently of the lunar
module. This communications equipment is capable of relaying voice and
telemetry data from the moon to the earth as well as transmitting color
television pictures. The equipment also provides the capability for reception
of voice communications from the earth, relay of voice to the crew, and
ground-command control of the television camera. The lunar roving vehicle
radio equipment, technically called the lunar communications relay unit,
employs a VHF radio link between the lunar roving vehicle and earth. The
color television camera with its positioning assembly, technically called the
ground commanded television assembly, is connected to the lunar
communications relay unit by a cable which carries ground commands to the
television control unit and returns the television pictures to the lunar
communications relay unit for transmission to earth. The crewmen communicate
directly with each other using their extravehicular communications systems.
Three batteries per crewman are provided for the three traverses. However, a
connection is made to the lunar roving vehicle power system when the
communications equipment is placed on the vehicle to provide a backup power
system for communications. A functional diagram of the lunar communications
relay unit is shown in
The lunar communications relay unit, and its S-band high-gain antenna
are installed on the forward chassis of the lunar roving vehicle by the
crew after vehicle deployment on the lunar surface. The S-band low-gain
antenna is installed into the lunar roving vehicle left inboard handhold.
The lunar communications relay unit is thermally controlled by three means:
thermal blankets regulate the exposed radiating surface and insulate the
unit from external environment; secondary-surface radiating mirrors reflect
undesired solar heat and emit undesired heat generated within the lunar
communications relay unit; and change-of-phase wax packages absorb excess
heat and stabilize the unit temperature through an absorption-discharge
A.4 EXPERIMENT EQUIPMENT
A.4.1 Lunar Surface Science Equipment
Descriptions of all of the Apollo 15 lunar surface science equipment
may be found in previous Apollo mission reports (references 8 through 11);
therefore, descriptions are not repeated here.
the Apollo lunar surface experiment package, and
geological tools used on Apollo 15.
lists the lunar surface
experiments and identifies the previous missions on which similar
experiments were deployed or conducted.
A.4.2 Inflight Science Equipment
Twelve experiments and several photographic activities utilized
equipment aboard the command and service modules during flight. Standard
spacecraft equipment was used to perform some scientific tasks. However, most
inflight science equipment was located in the scientific instrument module in
sector I of the service module. A view of the equipment in the scientific
instrument module, including some camera equipment, is shown in figure A-1.
All other cameras that were used for inflight experiments or photography were
located in the command module. The equipment used and the kinds of
information desired from each experiment and photographic activity are
described in the following paragraphs.
Gamma-ray spectrometer.- The gamma ray spectrometer experiment (S-160)
was conducted while in lunar orbit to obtain data on the degree of chemical
differentiation that the moon has undergone and the composition of the lunar
surface. The equipment was also operated during transearth coast to provide
calibration data on spacecraft and space background fluxes, and provide data
on galactic gamma-ray flux. A gamma-ray detector, capable of measuring gamma
radiation in the energy range from 200 000 to 10 million electron volts, was
mounted on a 25-foot boom located in the scientific instrument module (fig. A-
1). The boom could be fully extended or extended to two intermediate
positions, retracted, or jettisoned by the crew using controls in the command
module crew station. Controls were also provided to activate or deactivate
the spectrometer, incrementally alter the sensitivity (gain) of the detector,
and select either of two detector counting modes.
X-Rav fluorescence.- The X-ray fluorescence experiment (S-161)
equipment consisted of an X-ray detector assembly capable of detecting X-rays
the energy range from 1000 to 6000 electron volts, a solar monitor, and an X-
ray processor assembly. The X-ray detector assembly, located in the
scientific instrument module (fig. A-1), detected X-rays reflected from the
moon's surface or emitted by galactic X-ray sources. The solar monitor,
mounted in sector IV of the service module (displaced 1800 from the X-ray
detector assembly), measured solar X-ray flux. The measurement of fluorescent X-
ray flux from the lunar surface and the direct solar X-ray flux which
produces the fluorescence was expected to yield information on the nature of
the lunar surface material and the homogeneity of the upper few millimeters
of the lunar surface. Deep space measurements were expected to provide
information on galactic X-ray sources. Controls were provided in the command
module crew station to activate and deactivate the experiment, open the solar
monitor door, and open and close the X-ray detector protective cover.
Alpha particle spectrometer.- The alpha particle spectrometer
experiment (S-162) was designed to gather data to be considered along with
the gamma-ray and X-ray data in mapping the lunar chemical composition. The
types of information desired from this experiment were the gross rate of
lunar surface radon evolution and localized sources of enhanced radon
emission. In addition, transearth coast data were desired for background
and engineering evaluation of the alpha-particle and X-ray spectrometers.
The experiment equipment., consisted of an alpha particle sensing assembly
which could detect alpha particles in the energy range from 3.5 million to 7.5
million electron volts, supporting electronics, and temperature monitors
housed in the same enclosure as the X-ray fluorescence experiment assembly
(fig. A-1). Controls were provided in the command module crew station to
deploy a shield protecting the experiment detectors from spacecraft
contamination sources, and to activate and deactivate the experiment.
Mass spectrometer.- The mass spectrometer experiment (S-165) was conducted
to obtain data on the composition of the lunar ambient atmosphere as an aid
in understanding the mechanisms of release of gases from the surface, as a
tool to locate areas of volcanism, and as a means of determining the
distribution of gases in the lunar atmosphere. The experiment assembly
consisted of the mass spectrometer and its electronic components mounted on
a 24-foot boom which was extended from the scientific instrument module
(fig. A-1). The instrument was capable of measuring the abundance of
particles in the 12- to 66-atomic-mass-unit range. A shelf-mounted shield to
protect the spectrometer from spacecraft contamination sources when in its
stowed position opened and closed automatically when the boom was extended
and retracted. In addition to acquiring data while in lunar orbit, the
spectrometer was to be operated at various intermediate boom positions for
specified periods during transearth coast to determine the concentration of
constituents forming the command and service module contamination "cloud."
Command module crew station controls were provided to extend, retract, and
jettison the boom; activate/deactivate the spectrometer; select high and low
spectrometer discrimination modes, and multiplier gains; and
activate/deactivate the spectrometer ion source heaters and filaments.
S-band transponder (command and service module/lunar module).- The
command and service module and/or lunar module were tracked in lunar
orbit using the S-band transponders and high-gain antenna that were
normal vehicle equipment. The S-band Doppler resolver tracking data
obtained will be used to help determine the distribution of mass along
the lunar ground track. Tracking data were to be obtained from the docked
command and service module/lunar module while in the 170- by 60-mile
elliptical orbit, the 60-mile circular orbit, and the low-altitude
portion of the 60- by 8-mile elliptical orbit. Data were also to be
obtained from the undocked command and service module during the
unpowered portions of the 60-mile circular orbit, and from the undocked
lunar module during unpowered portions of flight.
Subsatellite experiments.- The subsatellite, launched from the command and service module
during lunar orbit, is the host carrier for three experiments for which data will be acquired over a
planned one-year period. The experiments are:
a. S-band transponder (S-164)
b. Particle shadows/boundary layer (S-173)
c. Subsatellite magnetometer (S-174)
The basic elements of the system, in addition to the subsatellite
itself, consisted of a mechanism
to deploy and launch the subsatellite from the scientific instrument
module, and a housing which
encased the subsatellite and its deployment/launcher device
The subsatellite contains charged particle telescope detectors
capable of detecting electrons in the
energy range from 20 000 to 320 000 electron volts and protons
in the energy range from 50 000 to 2.3
million electron volts. Spherical electrostatic analyzer detectors
are used to detect electrons in selected
energy bands from 580 to 15 000 electron volts. In addition,
the subsatellite contains a biaxial fluxgate
magnetometer which acquires data over a dynamic range of
±200 gammas, an optical solar aspect system for attitude
determination, a data storage unit, an S-band communications
system, and a power system. The primary power source consists
of solar cells on the subsatellite external surfaces. A rechargeable
silver cadmium battery is the secondary source of power that
sustains operation during passage of the
subsatellite through shadow. The subsatellite is hexagonal in
shape, 30 inches in length, and weighs
approximately 85 pounds. It has three equally-spaced booms
mounted around its base that deployed
automatically at launch to a length of 5 feet. The magnetometer
is mounted at the end of one boom,
whereas, the only purpose of the other two booms is to achieve
the desired spin-stabilization
characteristics. The subsatellite is shown in
Controls in the command module crew station were used for
launching the subsatellite and
retracting the deployment/launcher mechanism. The relative
parting velocity was approximately 4
ft/sec and the subsatellite was spin-stabilized at approximately
12 revolutions per minute about a spin
axis nearly perpendicular to the ecliptic plane.
S-band transponder experiment: Two-way S-band Doppler tracking
measurements of the subsatellite are made to obtain lunar
gravitational field data in addition to the data obtained
from tracking the command and service module and the lunar
Particle shadows/boundary layer experiment: The charged
particle detectors, the electrostatic
analyzer detectors, and the subsatellite support systems
are used to obtain data to study the formation
and dynamics of the earth's magnetosphere, the interaction
of plasmas with the moon, and the
physics of solar flares.
Subsatellite magnetometer: The magnetometer and the subsatellite
support systems are used to make magnetic field measurements in
lunar orbit. These data will be used in studies of the physical
properties of the moon and the interaction of plasmas with the moon.
Bistatic radar.- This experiment, technically designated "Downlink
Bistatic Radar Observations of the Moon" (S-170), was conducted to provide
fundamental new information on the upper few meters of the lunar crust,
and to provide engineering and calibration data needed for similar
experiments planned for the future. While the command and service
module was in
lunar orbit, S-band and VHF signals were transmitted from the
spacecraft, reflected from the lunar surface, and recorded on
the earth for subsequent
analysis. The high-gain antenna was preferred for S-band, although an
omnidirectional antenna was acceptable. The scimitar antenna was used for
VHF. The crew was required to maintain an attitude in
which the antenna was
pointed toward the lunar surface during the time that bistatic radar
measurements were being made.
Ultraviolet Photography - Earth and Moon.- Ultraviolet photography (S-
177) of the earth was obtained from earth orbit, from different points during
translunar and transearth coast, and from lunar orbit to determine
ultraviolet emission characteristics of the earth's atmosphere. A portion of
the photographs taken from lunar orbit were of the lunar surface. These will
be used to extend the wavelength range of ground-based colormetric work and
search for short-wavelength fluorescence.
The photographs were taken with a 70-mm Hasselblad electric camera and
105-mm ultraviolet transmitting lens. The camera was mounted on a bracket in
the right-hand side window. Two ultraviolet band-pass filters (centered at
3750 and 2600 angstrom) and a visual-range filter (4000 to 6000 angstrom)
were used. For each sequence of photographs requested, a minimum of four
were to be taken using black-and-white film and the aforementioned
filters, while one was to be taken using color film and a visual
range filter. The crew was required to install the mounting bracket,
mount and operate the camera, attach the filter slide and lens, change
filters, and record the exposure time. The crew was also required to maintain
the proper spacecraft attitude and attitude rates for each sequence.
Gegenschein from lunar orbit..- The Gegenschein from lunar orbit
experiment (S-178) -required three sequences of photographs to
be taken from the command module while in the shadow of the
moon - one in the direction of the antisolar vector, one in
the direction of the Moulton point, and one midway between
these two. A Nikon 35-mm camera and 55-mm lens were
used to obtain the photographs. The camera was mounted in
rendezvous window on a fixed mounting bracket. Window shades
and a darkened spacecraft were required to minimize the
effects of stray light from the spacecraft. The crew was
required to maneuver the spacecraft to the proper attitude
(the mission control center provided the proper spacecraft
orientation for camera pointing), inhibit the reaction
control system engines after spacecraft attitude rates
had been damped, operate the camera, and record the exposure time.
Apollo window meteroid.- The Apollo window meteroid experiment (S-176)
utilizes the command module windows as meteroid detectors and collectors.
Data are obtained by high-magnification scanning of the windows before and
after the flight.
Service module orbital photographic tasks.- These photographic tasks
comprised a detailed objective which required the use of the 24-inch
panoramic camera assembly, the 3-inch mapping camera assembly, and the
laser altimeter, all mounted in the scientific instrument module (fig. A-1).
Twenty-four-inch panoramic camera: This camera was included to obtain
high-resolution (1- to 2-meters from an altitude of 60 miles) panoramic
photographs with stereoscopic and monoscopic coverage of the lunar
surface. The photographs will aid in the correlation of other orbital
science data. The camera assembly consisted of a roll frame assembly, a
gimbal assembly to provide stereo coverage and forward motion
compensation, a main frame, a gaseous nitrogen pressure vessel to provide
gas for certain bearings, an optics system, a film drive and control
system, and a film cassette (that was required to be retrieved by an
extravehicular crewman during transearth coast). The camera did not
require deployment for operation. Controls were provided in the crew
station to activate/deactivate camera heaters, supply/remove primary
camera power, select operate/standby operation modes, supply film roller
torque to prevent slack in film during launch and maneuvers, activate a
five-frame film advance cycle if the camera was not operated in a 24-hour
period, increase/decrease the width of the exposure slit, and select the
stereo or monoscopic mode of operation.
Three-inch mapping camera: This camera was provided to obtain
high quality metric photographs of the lunar surface and stellar
photographs exposed simultaneously with the metric photographs. The lunar
surface photographs will aid in the correlation of experiment data with
lunar surface features. The stellar photographs provide a reference to
determine the laser altimeter pointing vector and the cartographic lens
pointing vector. The resolution capability of the metric camera was
approximately 20 meters from a distance of 60 miles. The metric and stellar
camera subsystems were integrated into a single unit which was deployed on
a rail-type mechanism in order to provide an unobstructed field of view
for the stellar camera. The system used the same gaseous nitrogen source
as the panoramic camera to provide an inert pressurized atmosphere within
the cameras to minimize potential static electrical corona discharge which
could expose film areas. In addition to the optics, the camera system
included a film drive/exposure/takeup system and a removable cassette
(that was required to be retrieved by an extravehicular crewman during
transearth coast). Controls were provided in the crew station to
activate/deactivate camera heaters and functions, compensate for image
motion and extend/retract the camera on its deployment rails.
Laser altimeter: The laser altimeter was furnished to obtain data on
the altitude of the command and service module above the lunar surface.
These data, acquired with a 1-meter resolution, were to support mapping and
panoramic camera photography as well as other lunar orbital experiments.
The laser altimeter could operate in either of two modes:
a. When the mapping camera was operating, the altimeter
automatically emitted a laser pulse to correspond to mid-frame
ranging for each
film frame exposed.
Command module controls were provided to activate/deactivate the
b. The altimeter could be decoupled from the mapping camera
to allow independent ranging measurements (one every 20 seconds).
Command module photographic tasks.- Photographs were to be obtained of:
a. Lunar surface areas of high scientific interest and of specific
portions of the lunar surface near the terminator.
These tasks involved the use of the following operational cameras:
b. Diffused galactic light of celestial objects, solar corona, the
lunar libration region, and the zodiacal light.
c. The lunar surface to extend selenodetic control and mapping.
d. The moon during lunar eclipse by the earth, and of a comet if
appropriate trajectory and celestial conditions existed.
a. A 16-mm data acquisition camera with an 18-mm lens.
Crew participation was required to operate the cameras, change lenses
and camera settings, record identification data, control the spacecraft
attitude and attitude rates, and control cabin illumination.
b. A 70-mm Hasselblad electric camera with 80-mm and 250-mm 'Lenses.
c . A 35-mm camera with a 55-mm lens.
A.5 SUMMARY OF PHOTOGRAPHIC EQUIPMENT
Nearly all experiments and detailed objectives require photography
either as a primary data source or for validation purposes. Photographic
equipment required for acquisition of data for experiments has
been discussed in conjunction with the applicable experiments in the preceding
section. For convenience, this equipment is also summarized in
along with photographic equipment required for other activities.
A.6 MASS PROPERTIES
Mass properties for the Apollo 15 mission are summarized in
These data represent the conditions as determined from postflight analyses of
expendable loadings and usage during the flight. Variations in command and
service modules and lunar module mass properties are determined for each
significant mission phase from lift-off through landing. Expendables usage are
based on reported real-time and postflight data as presented in other sections
of this report. The weights and center-of-gravity of the individual modules
(command, service, ascent stage, and descent stage) were measured prior to
flight and inertia values calculated. All changes incorporated after the
actual weighing were monitored, and the mass properties were updated.