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Command/Service module
The Command/Service Module (CSM) was one of two spacecraft that were utilized for the United States Apollo program, along with the Lunar Module, to land astronauts on the Moon. It was built for NASA by North American Aviation.
The spacecraft consisted of two segments: the Command Module, a (reentry capsule) which housed the crew and the equipment needed for re-entry and splashdown, and a Service Module that provided propulsion, electrical power and storage for various consumables required during a mission. The service module would be cast off and allowed to burn up in the atmosphere before the command module re-entered and brought the crew home.
After the conclusion of the Apollo program, the CSM saw service as a crew shuttle for the Skylab program, and also in the Apollo-Soyuz Test Project where a CSM docked with a Soviet Soyuz spacecraft in Earth orbit.
In 1964, NASA decided to build the Command/Service Module in two versions:
Block I spacecraft were used for all unmanned Saturn 1B and Saturn V test flights. Initially two manned flights were planned, but this was reduced to one in late 1966. This mission, designated AS-204 but named Apollo 1 by its flight crew, was planned for launch on February 21, 1967. But during a dress rehearsal for the launch on January 27, all three astronauts (Virgil I. "Gus" Grissom, Edward H. White, II and Roger Chaffee), were killed in a cabin fire which revealed serious design, construction and maintenance shortcomings in Block I, many of which were carried over into Block II.
After a thorough investigation by the Apollo 204 Review Board, it was decided to terminate the manned Block I phase and redefine Block II to incorporate the review board's recommendations. Block II incorporated a revised CM heat shield design, which was tested on the unmanned Apollo 4 and Apollo 6 flights, so the first all-up Block II spacecraft flew on the first manned mission, Apollo 7.
The two blocks were essentially similar in overall dimensions, but several design improvements resulted in weight reduction in Block II. The Apollo 1 spacecraft weighed 45,000 lb (20,412 kg), while Apollo 7 weighed only 36,993 lb. (16,520 kg.) Also, the Block I SM propellant tanks were slightly larger than in Block II. In the specifications given below, unless otherwise noted, all weights given are for the Block II spacecraft.
The Command Module was a truncated cone (frustum) measuring 10 feet 7 inches (3.2 m) tall and having a diameter of 12 feet 10 inches (3.9 m) across the base. The forward compartment contained two reaction control engines, the docking tunnel, and the components of the Earth Landing System. The inner pressure vessel housed the crew accommodations, equipment bays, controls and displays, and many spacecraft systems. The last section, the aft compartment, contained 10 reaction control engines and their related propellant tanks, fresh water tanks, and the CSM umbilical cables.
The command module's inner structure was an aluminum "sandwich" consisting of a welded aluminum inner skin, a thermally bonded honeycomb core, and a thin aluminum "face sheet". The central heat shield consisted of 40 individual panels interspersed with several holes and openings for the reaction control engines and after-compartment equipment access. The central compartment structure consisted of an inner aluminum face sheet with a steel honeycomb core, a glass-phenolic ablative honeycomb heat shield, a layer of q-felt fibrous insulation, a pore seal, a moisture barrier, and a layer of aluminized PET film thermal strips.
The aft heat shield consisted of four brazed honeycomb panels, four spot-welded sheet metal fairings, and a circumferential ring. The fairing segments were attached to the honeycomb panels and ring with conventional fasteners. The steel honeycomb core and outer face sheets were then thermally bonded to the inner skin in a giant autoclave. The aft heat shield is nearly identical to the central, with the exception of the outer alluminized PET film layer.
The components of the ELS were housed around the forward docking tunnel. The forward compartment was separated from the central by a bulkhead and was divided into four 90-degree wedges. The ELS consists of three main parachutes, three pilot parachutes, two drogue parachute motors, three upright bags, a sea recovery cable, a dye marker, and a swimmer umbilical.
The CM's center of mass was offset a foot or so from the center of pressure (along the symmetry axis). This provided a rotational moment during reentry, angling the capsule and providing some lift (a lift to drag ratio of about 0.368). The capsule was then steered by rotating the capsule using thrusters; when no steering was required, the capsule was spun slowly, and the lift effects cancelled out. This system greatly reduced the g-force experienced by the astronauts, permitted a reasonable amount of cross range and allowed the capsule to be targeted within a few miles.
At 24,000 feet (7.3 km) the forward heat shield was jettisoned using four pressurized-gas compression springs. The drogue parachutes were then released and slowed the spacecraft to 125 miles per hour (201 km/h). At 10,700 feet (3.3 km) the drogues were jettisoned. The pilot parachutes were deployed, which pulled out the mains. These slowed the CM to 22 miles per hour (35 km/h) for landing. The portion of the CM which made first contact with the water surface was built with crushable ribs to further mitigate the impact. The Apollo CM could safely parachute to an ocean landing with at least two parachutes (as it happened on Apollo 15), as the third parachute acted as a safety precaution.
The Command Module attitude control system consisted of twelve 93-pound-force (410 N) attitude control jets; ten were located in the aft compartment, and two pitch motors in the forward compartment. Four tanks stored 270 pounds (120 kg) of mono-methyl hydrazine fuel and nitrogen tetroxide oxidizer. They were pressurized by 1.1 pounds (0.50 kg) of helium stored at 4,150 pounds per square inch (28.6 MPa) in two tanks.
The forward docking hatch was mounted at the top of the docking tunnel. It was 30 inches (0.76 m) in diameter and weighed 80 pounds (36 kg). It was constructed from two machined rings that were weld-joined to a brazed honeycomb panel. The exterior side was covered with a 0.5-inch (13 mm) of insulation and a layer of aluminum foil. It was latched in six places and operated by a pump handle. At the center was a pressure equalization valve, used to equalize the pressure in the tunnel and lunar module before the hatch was removed.
The Unified Crew Hatch (UCH) measured 29 inches (74 cm) high, 34 inches (864 mm) wide, and weighed 225 pounds (102 kg). It was operated by a pump handle, which drove a ratchet mechanism to open or close fifteen latches simultaneously.
The docking probe allowed the command and lunar modules to connect together and maintain a pressure-tight seal during the mission. It was mounted to the docking tunnel in three places, and was designed to fold easily for storage.
The docking probe consisted of three shock attenuators, three tension linkages, a retractable and extendable probe, and two power umbilicals. Before docking, a crewman in the CM extended the probe. When it came in contact with the drogue, three capture latches in the probe head held the two modules together. The probe was retracted, pulling the CM and LM together. Twelve latches attached to the forward docking ring automatically activated to form an air-tight seal.
At LM separation, the probe and forward docking ring were pyrotechnically separated leaving all docking equipment with the lunar module.
The central pressure vessel of the command module was its sole habitable compartment. It had an interior volume of 210 cubic feet (5.9 m3) and housed the main control panels, crew seats, guidance and navigation systems, food and equipment lockers, the waste management system, and the docking tunnel.
Dominating the forward section of the cabin was the crescent-shaped main display panel measuring nearly seven feet (2.1 m) wide and three feet (0.9 m) tall. It was arranged into three panels, each emphasizing the duties of each crew member. The mission commander’s panel (left side) included the velocity, attitude, and altitude indicators, the primary flight controls, and the main FDAI (Flight Director Attitude Indicator).
The CM pilot served as navigator, so his control panel (center) included the Guidance and Navigation computer controls, the caution and warning indicator panel, the event timer, the Service Propulsion System and RCS controls, and the environmental control system controls.
The LM pilot served as systems engineer, so his control panel (right-hand side) included the fuel cell gauges and controls, the electrical and battery controls, and the communications controls.
Flanking the sides of the main panel were sets of smaller control panels. On the left side were a circuit breaker panel, audio controls, and the SCS power controls. On the right were additional circuit breakers and a redundant audio control panel, along with the environmental control switches. In total, the command module panels included 24 instruments, 566 switches, 40 event indicators, and 71 lights.
The three crew couches were constructed from hollow steel tubing and covered in a heavy, fireproof cloth known as Armalon. The leg pans of the two outer couches could be folded in a variety of positions, while the hip pan of the center couch could be disconnected and laid on the aft bulkhead. One rotation and one translation hand controller was installed on the armrests of the left-hand couch. The translation controller was used by the crew member performing the LM docking maneuver, usually the CM Pilot. The center and right-hand couches had duplicate rotational controllers. The couches were supported by eight shock-attenuating struts, designed to ease the impact of touchdown on water or, in case of an emergency landing, on solid ground.
The contiguous cabin space was organized into six equipment bays:
The CM had five windows. The two side windows measured 13 inches (330 mm) square next to the left and right-hand couches. Two forward-facing triangular rendezvous windows measured 8 by 13 inches (204 by 330 mm), used to aid in rendezvous and docking with the LM. The hatch window was 10 5/8 in. diameter (27cm) and was directly over the center couch. Each window assembly consisted of three thick panes of glass. The inner two panes, which were made of aluminosilicate, made up part of the module's pressure vessel. The fused silica outer pane served as both a debris shield and as part of the heat shield. Each pane had an anti-reflective coating and a blue-red reflective coating on the inner surface.
Specifications
The Service Module was an unpressurized cylindrical structure, measuring 24 feet 7 inches (7.5 m) long and 12 feet 10 inches (3.9 m) in diameter. The interior was a simple structure consisting of a central tunnel section 44 inches (1.1 m) in diameter, surrounded by six pie-shaped sectors. The sectors were topped by a forward bulkhead and fairing, separated by six radial beams, covered on the outside by four honeycomb panels, and supported by an aft bulkhead and engine heat shield. The sectors were not all equal 60° angles, but varied according to required size.
The forward fairing measured 2 feet 10 inches (864 mm) long and included the Reaction Control System (RCS) computer, umbilical connection, power distribution block, ECS controller, separation controller, components for the high-gain antenna, and eight EPS radiators. The umbilical housing contained the main electrical and plumbing connections to the CM. The fairing externally contained a retractable forward-facing spotlight; an EVA floodlight to aid the Command Module pilot in SIM film retrieval; and a flashing rendezvous beacon visible from 54 nautical miles (100 km) away as a navigation aid for rendezvous with the Lunar Module (LM).
The SM was connected to the CM using three tension ties and six compression pads. The tension ties were stainless steel straps bolted to the CM's aft heat shield. It remained attached to the Command Module throughout most of the mission, until being jettisoned just prior to re-entry into the Earth's atmosphere. At jettison, the CM umbillical connections were cut using a pyrotechnic-activated guillotine assembly. Following jettison, the SM aft translation thrusters automatically fired continuously to distance it from the CM, until either the RCS fuel or the fuel cell power was depleted. The roll thrusters were also fired for five seconds to make sure it followed a different trajectory from the CM and faster break-up on re-entry.
The 20,500-pound-force (91,000 N) SPS engine was used to place the Apollo spacecraft into and out of lunar orbit, and for mid-course corrections between the Earth and Moon. The engine used was an AJ10-137 engine using Aerozine 50 as fuel and nitrogen tetroxide (N2O4) as oxidizer. The propellants were pressure-fed to the engine by 39.2 cubic feet (1.11 m3) of gaseous helium at 3600 psia, carried in two 40-inch (1.0 m) diameter spherical tanks.
The engine measured 152.82 inches (3.882 m) long and 98.48 inches (2.501 m) wide at the base. It was mounted on two gimbals to provide pitch and yaw control in lieu of the RCS during SPS firings. The combustion chamber and pressurant tanks were housed in the central tunnel.
The thrust level was actually twice what was needed to accomplish the lunar orbit rendezvous (LOR) mission mode, because the engine was originally sized to lift the CM with a much larger SM off of the lunar surface in the direct ascent mode assumed in original planning (see Choosing a mission mode.) A contract was signed in April 1962 for the Aerojet-General company to start developing the engine, before the LOR mode was officially chosen in July of that year.
Four clusters of four reaction control system (RCS) thrusters were installed around the upper section of the SM every 90°. The sixteen-thruster arrangement provided rotation and translation control in all three spacecraft axes. Each thruster generated 100 pounds (45 N) of thrust, and used mono-methyl hydrazine (MMH) as fuel and nitrogen tetroxide as oxidizer. Each quad assembly measured 8 feet (2.4 m) by 3 feet (0.91 m) and had its own fuel tank, oxidizer tank, helium pressurant tank, and associated valves and regulators.
The Lunar Module used a similar four-quad arrangement of the identical thruster engines for its RCS.
Electrical power was produced by three fuel cells, each measuring 44 inches (1.1 m) tall by 22 inches (0.56 m) in diameter and weighing 245 pounds (111 kg). These combined hydrogen and oxygen to generate electrical power, along with some of the water used for drinking and other purposes. The cells were fed by two hemispherical-cyllindrical 31.75-inch (0.806 m) diameter tanks, each holding 29 pounds (13 kg) of liquid hydrogen, and two spherical 26-inch (0.66 m) diameter tanks, each holding 326 pounds (148 kg) of liquid oxygen (which also supplied the environmental control system).
On the flight of Apollo 13, the EPS was disabled by an explosive rupture of one oxygen tank, which punctured the second tank and led to the loss of all oxygen. After the accident, a third oxygen tank was added to prevent operation below 50% tank capacity which allowed removal of the tank's internal stirring fan equipment, which had contributed to the failure.
Storage tanks were carried for water and oxygen. Waste heat from the CM cabin was dumped to space by two 30-square-foot (2.8 m2) radiators located on the lower section of the exterior walls, one covering sectors 2 and 3, and the other covering sectors 5 and 6.
Short-range communications with the Lunar Module employed two VHF scimitar antennas mounted on the exterior walls, just above the ECS radiators.
A steerable S-band high-gain antenna, used for earth-to-moon communications, was mounted on the aft bulkhead. This was an array of four 31-inch (0.79 m) diameter reflectors surrounding a single 11-inch (0.28 m) square reflector. During launch it was folded down parallel to the main engine to fit inside the Spacecraft-to-LM Adapter (SLA). After CSM separation from the SLA, it deployed at a right angle to the SM.
The Low Earth Orbit payload capability of the Saturn IB booster used to launch the Low Earth Orbit missions (Apollo 1 (planned), Apollo 7, Skylab 2, Skylab 3, Skylab 4, and Apollo-Soyuz) could not handle the 66,900 pounds (30,300 kg) mass of the fully fueled CSM. This was not a problem because the delta-V requirement of these missions was much smaller than that of the lunar mission, therefore they were launched with less than half of the full SPS propellant load. The CSMs launched in orbit on Saturn IB ranged from 32,558 pounds (14,768 kg) (Apollo-Soyuz), to 46,000 pounds (21,000 kg) (Skylab 4).
Since the Saturn IB manned missions didn't require communication at lunar distances, the high-gain antenna was not required; the VHF antennas sufficed for communication with the ground. The S-band antenna was omitted from Apollo 1, Apollo 7, and the three Skylab flights, though included on the Apollo-Soyuz mission.
On the Skylab and Apollo-Soyuz missions, some additional dry weight was saved by removing the otherwise empty fuel and oxidizer storage tanks (leaving the partially filled sump tanks), along with one of the two helium pressurant tanks. This permitted the addition of some extra RCS propellant to allow for use as a backup for the deorbit burn in case of possible SPS failure.
Since the spacecraft for the Skylab missions would not be occupied for most of the mission, there was lower demand on the power system and one of the three fuel cells was deleted from these SMs.
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