Project Apollo: The Beginnings | Mission Planning | Landing Site Selection | Earthbound Support Systems
Astronaut Selection and Training | The Saturn V | The Saturn 1B | The Apollo Spacecraft
Guidance and Navigation | Command and Service Modules | The Lunar Module
Assembling and Launching | Pathfinders | The Early Missions | Apollo 11, The First Landing
The Intermediate Missions | Apollo 15 Exploration | Apollo 16 Exploration | Apollo 17 Exploration
Skylab and Apollo-Soyuz | Conclusion

The CSM was designed and built by North American Aviation at its Downey, California plant. The initial contract required the delivery of 11 non-flying mock-up models for configuration analysis, 15 boilerplate test vehicles, and 11 flight spacecraft. The flight vehicles evolved into Block I, those built without a forward docking facility for use in earth orbit for testing and flight operations training, and Block II, those with the docking equipment for full lunar operations.

By early in 1964, the concept of In Flight Maintenance (IFM), in which it was envisaged that the crew would carry out repairs to malfunctioning equipment during flight, was abandoned. Carrying spares and tools involved too high a weight penalty and experience had showed that events during flight emergencies would probably occur and develop into an unmanageable situation too rapidly for repairs by the crew to be carried out. Instead, NASA required the main contractors to design in reliability and switchable redundancy whereby malfunctioning components could be by-passed or isolated during flight and a secondary system used instead.

Also by 1964, NAA studies showed that the problems inherent in recovery of the spacecraft over land were very nearly insurmountable. Impacting the ground without the use of retro rockets was almost certainly going to injure the crew. In April 1964 the decision to recover the command module by parachute with a splashdown in the sea and recovery by US naval vessels was accepted and approved by NASA.

Difficulties and delays were experienced by both NAA and Grumman with the ever changing requirements of the mission concept as it evolved. NAA design for the Block II spacecraft was being hindered by Grumman's delay in settling on a final design for the LM and in particular for the docking arrangement between the two spacecraft. For the first four months of 1964 Grumman led a mission study together with NASA, NAA and MIT to investigate the exact requirements of the hardware from lift off to splashdown. Using an arbitrary date of 6 May 1968 for the first manned landing and the basic premise to '...land two men on the moon's surface, carrying 250 pounds of scientific equipment and return them with 100 pounds of samples', they were able to calculate the exact trajectories, fuel loads and other criteria necessary to finalise the hardware design. This produced the Design Reference Mission (DRM), a three volume document specifying the exact requirements by which contractors would design the hardware.

Unresolved problems with the design of both Block I and Block II spacecraft brought about a change in North American's design team when Paup was replaced by Dale Myers as North American's Apollo program manager. NAA relationship with NASA became strained at times, especially during the first years of the spacecraft's development. Reports of poor workmanship and safety practices troubled NASA project officials and the situation came to a head in 1965 when NASA Apollo program manager, General Sam Phillips undertook a review and wrote to North American pointing out various shortcomings in their procedures. Improvements were made, but on 27 January, 1967, the first Block I, manned flight, Apollo spacecraft caught fire on its launchpad killing its crew Gus Grissom, Edward White and Roger Chaffee. The inquiry that followed was unable to pinpoint the exact cause of the fire but its report highlighted deficiencies in North American's design, workmanship and quality control.

The Apollo 1 fire forced North American to carry out a complete revision of the command module design removing as many sources of flammable material as possible and incorporating fireproofing features to electrical wiring and connections, all of which delayed the start of production. Harrison Storms stood down from the Apollo project, as did NASA's Joseph Shea who felt a personal responsibility for the disaster. Storms was replaced by William D Bergan and Shea by George Low from NASA headquarters, who took a demotion to step into the position of ASPO manager. Bergan replaced key North American staff and introduced a system of individual teams for each spacecraft. This practice was also adopted by the LM's manufacturer, Grumman who also were forced to review the LM's design.

Command Module (CM)

The CM was the real heart of the Apollo spacecraft as it contained the three man crew, navigation and environmental systems for the overall flight and was the only part of the complete stack to return intact to earth. It was conical in shape with a blunt, convex base, measuring 12 feet 10 inches in diameter and 11 feet high and weighed 12,500 pounds when unloaded

The body of the spacecraft was constructed in a double layer separated by a thermal insulation layer. An inner pressure shell was formed from a lightweight, double skinned, alloy matrix and an outer skin, made from a honeycomb steel alloy doubled up as a heat shield and micro-meteorite protective layer. The CM shell was separated into three main compartments, the forward, crew and aft compartments and the distribution of the craft's weight was carefully calculated to provide an off-set to its centre of gravity. This induced a natural stabilising force during re-entry and provided aerodynamic lift. Attitude control by the on-board computer to make adjustments to its flight trajectory was obtained through the RCS and made use of the craft's lifting body characteristics.

Two individual RCS systems were utilised by the CSM. The system utilised in the service module was manufactured by the Marquardt company of California, and incorporated small 100 pound thruster rockets, clustered in groups of four (quads), in two separate sub-systems which were spaced equidistant around the circumference of the craft. The command module used a separate system designed by Rocketdyne, which, of necessity, was flush fitting with the spacecraft's outer surface to remove protruding equipment that would be damaged by the aerodynamic forces generated during re-entry. The two systems, when used in various combinations with each other controlled the pitch, roll and yaw of the combined CSM and to make minor adjustments to the crafts speed. Both RCS systems used the same hypergolic fuels and pressurisation as the SPS engine.

Forward Compartment

The forward compartment surrounded the forward docking tunnel and contained the Earth Landing System (ELS). This equipment, for use during the earth re-entry phase, was protected by a heatshield, which was jettisoned after passing through the thermal interface of re-entry. The ELS provided the deployment of two drogue parachutes of 16.5 feet diameter by pyrotechnic mortars at about 24,000 feet altitude to stabilise and slow the latter part of the descent. At 10,000 feet the drogues would be discarded and three main descent parachutes deployed by mortars to complete the descent to splashdown.

Three inflatable bags were also housed within the forward compartment to right the spacecraft if it overturned at splashdown. Stable 1 was the desirable upright floating condition (apex up) and Stable 2 the inverted position. Two reaction control motors were also housed in the forward compartment to provide pitch attitude control at the forward end of the spacecraft.

Crew Compartment

The interior crew compartment provided a living space of 210 cubic feet, similar to the capacity of a large family saloon car and was the only part of the CM to be pressurised with an internal atmosphere. During take off the crew occupied three couches situated side by side and facing the main control and instrument panel. The main control panel carried the majority of the flight controls, switches, circuit breakers, warning lights and alarms required to control and monitor the spacecraft's performance during major manoeuvres. After take off the central couch could be collapsed to provide more room to move about the interior and gain access to the lower equipment bay situated centrally under the couches.

Two hatches allowed access to the crew compartment. The main side hatch, twenty nine by thirty four inches, was used for crew ingress and egress prior to and after the mission. The second, forward hatch, at the apex of the CM cone, allowed access to the thirty inch diameter docking tunnel, docking probe and drogue and to the lunar module when docked. Valves within the hatches enabled the crew to vent atmosphere to space or to equalise internal pressures between spacecraft.

The crew were provided with five windows in the CM. One nine inches diameter in the side hatch, Two windows, thirteen inches square either side of the hatch, used for observation and a further two triangular windows, angled to face forward, thirteen by eight inches, used during docking. All windows were constructed with triple panes with the outer pane nearly three-quarters of an inch thick. Each pane was treated to filter infra-red and ultra-violet light and could withstand temperatures of 2,800 degrees Fahrenheit.

Removal of the centre couch provided access to the lower equipment bay and the Guidance Navigation and Control System (GNCS) with its three main sub-systems, the Apollo Guidance Computer (AGC), the Inertial Measurement Unit (IMU) and the optical alignment system (OPS). The optical system was centrally placed under the centre couch position. The crew were provided with two DSKY's, one in the lower equipment bay for inputs from the OPS and a second housed in the main instrument panel.

Communication equipment, also housed adjacent to the AGC provided voice, television and telemetry communication with the Manned Space Flight Network (MSFN) and between command and lunar modules when separated during the landing and rendezvous manoeuvers. Communication with the MSFN was through a high gain, S-band antenna, mounted at the aft end of the service module and consisted of four, 31 inch diameter dishes mounted on a folding arm which was deployed after launch. Alignment of the antenna was controlled by the AGC through the communication system. The crews vital signs were monitored through the Biomed monitoring system. A harness with sensors taped to the astronauts skin supplied heart rate, respiration and EKG information via a continuously linked information channel.

The Environmental Control System (ECS) situated in the left hand equipment bay, below the left hand couch, monitored and controlled the CM's internal atmosphere, pressure and temperature and controlled the temperature regulation of internal electronic equipment. Manufactured by the Hamilton Standard Co, it also controlled water production through the service module's fuel cells for cooling electronic equipment and supplied hot and cold potable water for the crew's consumption. Up to launch the internal pressure was maintained just above atmospheric, with a mixture of oxygen and nitrogen. During launch this pressure was bled away to be replaced by an internal atmosphere of pure oxygen at 5 psi and maintained at 75 degrees Fahrenheit. The ECS also monitored the condition of the internal atmosphere and filtered out carbon dioxide from the crew's exhaled breath by the use of replaceable lithium dioxide filters.

Control switches and circuit breakers occupied the bulkheads to either side of the couches. Bays and cupboards around the walls of the crew compartment and under the couches also housed all the equipment and supplies that would be necessary for the needs and comfort of the crew. Lockers on each side of the lower bay housed food, clothing, cameras, medical kit, personal hygiene kit and survival gear. These bays also provided storage space for the Sample Return Containers (SRC) that would house the moon rock samples on the return journey.

Food packages consisted mainly of dehydrated food in bags that could be reactivated by the addition of hot or cold water from the ECS water spigot in the left lower equipment bay. Individual meals of over 70 different items were supplied for the mission duration, each identified for day and crew member. A further snack pantry was provided to supplement the meals and prevent the raiding of meal combinations for favoured items. Waste disposal of urine and water was through a vent to space in the right lower equipment bay, while solid waste products were stowed , in defecation bags, in one of the cupboard spaces after treatment.

Aft Compartment

The aft compartment and side walls, were divided radially into twenty four bays to house the various systems and consumable for the flight including ten Marquardt reaction control engines with their propellant tanks and helium tanks for pressurisation. Water tanks and five, silver zinc oxide batteries for power to supply the CM after jettisoning the SM were also stowed here. Three of the batteries powered to the CM's electronic equipment during re-entry and two initiated the pyrotechnics used for CM/SM separation and parachute deployment.

An external umbilical harness connected the service module to the aft compartment of the CM to provide services between the two during the mission. On jettisoning the SM just prior to re-entry, a sequencer deadfaced all electrical circuits between the two modules and closed oxygen, water and other supply lines. Separation of the umbilical was achieved by a pyrotechnic powered guillotine which cut all interconnecting links between the two modules and retracted the umbilical arm clear of the CM. Further explosive bolts severed three stainless steel straps which secured the two modules together.

The rear, blunt face of the spacecraft was covered by the all important, re-entry heat shield. Its function was to safeguard the crew from the 5000 degrees Fahrenheit temperature generated by friction through the earth's atmosphere during the Thermal Interface (TI). The Apollo heatshield was a new design differing from the previous Gemini type. Manufactured by the Avco Corporation of Massachusetts, the heatshield consisted of a cover varying in thickness from two and half inches at the centre, to half an inch at the outer periphery. It was constructed from a brazed stainless steel honeycomb matrix whose 400,000 cells were filled with a fibre glass and phenolic resin ablation material. During re-entry the ablation material heated and burned off carrying the heat away from the spacecraft.

Service Module (SM)

The Service Module (SM), also built by North American Aviation with the CM, provided the motive and electrical power, oxygen and communication facility with earth. Cylindrical in shape, it measured 12 feet 10 inches in diameter and 24 feet 7 inches long and was attached to the CM during the flight. Constructed from one inch thick alloy honeycomb panels and partitioned off by milled aluminium radial beams internally into six longitudinal bays around a central tubular core, it weighed over 51,000 pounds.

The main Service Propulsion System (SPS) engine was housed on a thrust structure at the rear of the central core and two helium tanks to pressurise the fuel system took up space at the forward end. Fuel tanks for the SPS engine took up four of the bays while oxygen and hydrogen tanks and three fuel cells to provide the spacecraft's main water and electricity supply occupied a fifth. Power cells developed and manufactured by the Pratt and Whitney Aircraft Division provided electricity by combining oxygen and hydrogen which also produced water as a by-product. The water was passed forward through the umbilical connection to the CM's environmental system to be used for cooling electronic systems and as potable water for the crew's consumption. The remaining sixth bay, the Scientific Instrument Module (SIM) bay, was left empty until later Apollo flights when it carried remote sensing equipment for lunar data collection from orbit.

The SPS, a restartable, constant thrust, rocket engine, manufactured by Aerojet General, provided 20,500 pounds of thrust and was gimballed for directional thrust to provide the orbital and speed changes and mid course corrections to the spacecraft's trajectory. The engine was to be used for all large velocity changes of the spacecraft after the S-IVB third stage had been jettisoned. This would include braking the craft into a lunar orbit and when the time came, pushing it out of that orbit again into the homeward trajectory. The engine's firing time and alignment was controlled by the Stabilisation and Control System (SCS), a subsystem of the GNCS in the command module.