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.1.1 Structure and Thermal Systems

A scientific instrument module was installed in sector I of the service module ( fig. A-2). 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 ( fig. A-2). Protective 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 A.4.2.

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 ( fig. A-3).

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.1.3 Instrumentation

A scientific data system was integrated with the existing telemetry system ( fig. A-4) 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- code-modulated data.

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 603 ( fig. A-5) was added to provide umbilical connections for extravehicular activity. Panel 604 ( fig. A-5) was added to provide an audio warning signal to the extravehicular crewman in the event of low suit pressure or low oxygen flow.

A.1-5 Propulsion

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 gage.
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 ( fig. A-6) 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 umbilical.

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.2.1 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:

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).

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 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.

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.

A-2.4 Propulsion

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 engines.

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-3.1 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:

The portable life support system was modified to extend the lunar surface stay time capability. There were four major changes:

A.3.2 Lunar Roving Vehicle

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 display panel.

The chassis ( fig. A-8) 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 controller.

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 remaining battery.

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 distance traveled, 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 figure A-9.

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 cycle.


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. Figure A-10 illustrates the Apollo lunar surface experiment package, and figure A-11 shows the geological tools used on Apollo 15. Table A-I 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:

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 (fig. A-1).

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 figure A-12.

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 module.

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 and electrical 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 the right-hand 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:

Command module controls were provided to activate/deactivate the altimeter.

Command module photographic tasks.- Photographs were to be obtained of:

These tasks involved the use of the following operational cameras:

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.


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 table A-II along with photographic equipment required for other activities.


Mass properties for the Apollo 15 mission are summarized in table A-III. 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.