Apollo Command/Service Module

Apollo 15 Command and Service Module in lunar orbit

Apollo-Soyuz Test Project

Lunar orbit

Apollo insignia

Apollo Block II CSM diagram

The Command/Service Module (CSM) was one of two spacecraft, along with the Lunar Module, used for the United States Apollo program which landed astronauts on the Moon. It was built for NASA by North American Aviation. It was launched by itself on three suborbital and low Earth orbit Apollo test missions using the Saturn IB launch vehicle. It was also launched twelve times on the larger Saturn V launch vehicle, both by itself and with the Lunar Module. It made a total of nine manned flights to the Moon boosted by the Saturn V.

After the Apollo lunar program, the CSM saw manned service as a crew shuttle for the Skylab program, and the Apollo-Soyuz Test Project (ASTP) in which an American crew rendezvoused and docked with a Soviet Soyuz spacecraft in Earth orbit. Ending with ASTP in 1975, it had flown six astronaut missions to low Earth orbit only.

The CSM consisted of two segments: the Command Module, a cabin that housed a crew of three and 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 was cast off and allowed to burn up in the atmosphere before the Command Module re-entered and brought the crew home.

The CSM was initially designed to return all three astronauts from the lunar surface on a direct-descent mission which would not use a separate Lunar Module, and thus had no provisions for docking with another spacecraft. This, plus other required design changes led to the decision to design two versions of the CSM: Block I was to be used for unmanned missions and a single manned Earth orbit flight (Apollo 1), while the more advanced Block II was designed for use with the Lunar Module. The Apollo 1 flight was cancelled after a cabin fire killed the entire crew and destroyed the Command Module during a launch rehearsal test. Corrections of the problems which caused the fire were applied to the Block II spacecraft, which was used for all manned missions.

## Development History

When NASA awarded the initial Apollo contract to North American Aviation on November 28, 1961, it was still assumed the lunar landing would be achieved by direct descent rather than by lunar orbit rendezvous. Therefore, design proceeded without a means of docking the Command Module to a Lunar Excursion Module (LEM). But the change to lunar orbit rendezvous, plus several technical obstacles encountered in some subsystems (such as environmental control), soon made it clear that substantial redesign would be required. In 1963, NASA decided the most efficient way to keep the program on track was to proceed with the development in two versions:

* Block I would continue the preliminary design, to be used for early low Earth orbit test flights only.
* Block II would be the lunar-capable version, including a docking hatch and incorporating weight reduction and lessons learned in Block I. Detailed design of the docking capability depended on design of the LEM, which was contracted to Grumman Aircraft Engineering.

By January 1964, North American started presenting Block II design details to NASA. 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 (Gus Grissom, Ed White and Roger Chaffee), were killed in a cabin fire which revealed serious design, construction and maintenance shortcomings in Block I, many of which would have been 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. Also, the Block I Service Module propellant tanks were slightly larger than in Block II. The Apollo 1 spacecraft weighed approximately 45,000 pounds (20,000 kg), while the Block II Apollo 7 weighed 36,400 lb (16,500 kg). (These two Earth orbital craft were lighter than the craft which later went to the Moon, as they carried propellant in only one set of tanks, and did not carry the high-gain S-band antenna.) In the specifications given below, unless otherwise noted, all weights given are for the Block II spacecraft.

The total cost of the CSM for development and the units produced was $36.9B in 2016 dollars, adjusted from a nominal total of $3.7B using the NASA New Start Inflation Indices.

## Command Module

The Command Module was a truncated cone (frustum) 10 feet 7 inches (3.23 m) tall with a diameter of 12 feet 10 inches (3.91 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.

### **Construction**

The Command Module (CM) consisted of two basic structures joined together: the inner structure (pressure shell) and the outer structure.

The inner structure was an aluminium sandwich construction which consisted of a welded aluminium inner skin, adhesively bonded aluminium honeycomb core, and outer face sheet. The thickness of the honeycomb varied from about 1.5 inches (3.8 cm) at the base, to about 0.25 inches (0.64 cm) at the forward access tunnel. This inner structure was the pressurized crew compartment.

The outer structure was made of stainless steel brazed honeycomb brazed between steel alloy face sheets. It varied in thickness from 0.5 inch to 2.5 inches. Part of the area between the inner and outer shells was filled with a layer of fiberglass insulation as additional heat protection.

#### **Thermal Protection (Heat Shield)**

An ablative heat shield on the outside of the CM protected the capsule from the heat of re-entry, which is sufficient to melt most metals. This heat shield was composed of phenolic epoxy resin, a type of reinforced plastic. During re-entry, this material charred and melted away, absorbing and carrying away the intense heat in the process. The heat shield has several outer coverings: a pore seal, a moisture barrier (a white reflective coating), and a silver Mylar thermal coating that looks like aluminium foil.

The heat shield varied in thickness from 2 inches (5.1 cm) in the aft portion (the base of the capsule, which faced forward during re-entry) to 0.5 inches (1.3 cm) in the crew compartment and forward portions. Total weight of the shield was about 3,000 pounds (1,400 kg).

#### **Forward Compartment**

The forward compartment was the area outside the inner pressure shell in the nose of the capsule, located around the forward docking tunnel and covered by the forward heat shield. The compartment was divided into four 90-degree segments which contained Earth Ianding equipment (all the parachutes, recovery antennas and beacon light, and sea recovery sling), two reaction control engines, and the forward heat shield release mechanism.

At about 25,000 feet (7,600 m) during re-entry, the forward heat shield was jettisoned to expose the Earth landing equipment and permit deployment of the parachutes.

#### **Aft Compartment**

The aft compartment was located around the periphery of the command module at its widest part, just forward of (above) the aft heat shield. The compartment was divided into 24 bays containing 10 reaction control engines; the fuel, oxidizer, and helium tanks for the CM reaction control subsystem; water tanks; the crushable ribs of the impact attenuation system; and a number of instruments. The CM-SM umbilical, the point where wiring and plumbing ran from one module to the other, was also in the aft compartment. The panels of the heat shield covering the aft compartment were removable for maintenance of the equipment before flight.

### **Earth Landing System**

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 consisted of two drogue parachutes with mortars, three main parachutes, three pilot parachutes to deploy the mains, three inflation bags for uprighting the capsule if necessary, a sea recovery cable, a dye marker, and a swimmer umbilical.

The Command Module's centre of mass was offset a foot or so from the centre of pressure (along the symmetry axis). This provided a rotational moment during re-entry, 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 directional control and allowed the capsule's splashdown point 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 deployed, slowing the spacecraft to 125 miles per hour (201 kilometres per hour). At 10,700 feet (3.3 km) the drogues were jettisoned and the pilot parachutes, which pulled out the mains, were deployed. These slowed the CM to 22 miles per hour (35 kilometres per hour) for splashdown. The portion of the capsule that first contacted the water surface contained four crushable ribs to further mitigate the force of impact. The Command Module could safely parachute to an ocean landing with only two parachutes deployed (as occurred on Apollo 15), the third parachute being a safety precaution.

### **Reaction Control System**

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 monomethylhydrazine 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.

### **Hatches**

The forward docking hatch was mounted at the top of the docking tunnel. It was 30 inches (76 cm) 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 aluminium foil. It was latched in six places and operated by a pump handle. The hatch contained a valve in its centre, used to equalize the pressure between the tunnel and the CM so the hatch could be removed.

The Unified Crew Hatch (UCH) measured 29 inches (74 cm) high, 34 inches (86 cm) 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.

### **Docking Assembly**

The Apollo spacecraft docking mechanism was a non-androgynous system, consisting of a probe located in the nose of the CSM, which connected to the drogue, a truncated cone located on the Lunar Module. The probe was extended like a scissor jack to capture the drogue on initial contact, known as soft docking. Then the probe was retracted to pull the vehicles together and establish a firm connection, known as "hard docking". The mechanism was specified by NASA to have the following functions:

* Allow the two vehicles to connect, and attenuate excess movement and energy caused by docking
* Align and centre the two vehicles and pull them together for capture
* Provide a rigid structural connection between both vehicles, and be capable of removal and re-installation by a single crewman
* Provide a means of remote separation of both vehicles for the return to Earth, using pyrotechnic fasteners at the circumference of the CSM docking collar
* Provide redundant power and logic circuits for all electrical and pyrotechnic components.

#### **Coupling**

The probe head located in the CSM was self-centring and gimbal-mounted to the probe piston. As the probe head engaged in the opening of the drogue socket, three spring-loaded latches depressed and engaged. These latches allowed a so-called 'soft dock' state and enabled the pitch and yaw movements in the two vehicles to subside. Excess movement in the vehicles during the 'hard dock' process could cause damage to the docking ring and put stress on the upper tunnel. A depressed locking trigger link at each latch allowed a spring-loaded spool to move forward, maintaining the toggle linkage in an over-centre locked position. In the upper end of the Lunar Module tunnel, the drogue, which was constructed of 1-inch-thick aluminium honeycomb core, bonded front and back to aluminium face sheets, was the receiving end of the probe head capture latches.

#### **Retraction**

After the initial capture and stabilization of the vehicles, the probe was capable of exerting a closing force of 1,000 pounds-force (4.4 kN) to draw the vehicles together. This force was generated by gas pressure acting on the centre piston within the probe cylinder. Piston retraction compressed the probe and interface seals and actuated the 12 automatic ring latches which were located radially around the inner surface of the CSM docking ring. The latches were manually re-cocked in the docking tunnel by an astronaut after each hard docking event (lunar missions required two dockings).

#### **Separation**

An automatic extension latch attached to the probe cylinder body engaged and retained the probe centre piston in the retracted position. Before vehicle separation in lunar orbit, manual cocking of the twelve ring latches was accomplished. The separating force from the internal pressure in the tunnel area was then transmitted from the ring latches to the probe and drogue. In undocking, the release of the capture latches was accomplished by electrically energizing tandem-mounted DC rotary solenoids located in the centre piston. In a temperature degraded condition, a single motor release operation was done manually in the Lunar Module by depressing the locking spool through an open hole in the probe heads, while release from the CSM was done by rotating a release handle at the back of the probe to rotate the motor torque shaft manually. When the Command and Lunar Modules separated for the last time just before re-entry, the probe and forward docking ring were pyrotechnically separated, leaving all docking equipment attached to the lunar module. In the event of an abort during launch from Earth, the same system would have explosively jettisoned the docking ring and probe from the CM as it separated from the boost protective cover.

### **Cabin Interior Arrangement**

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 7 feet (2.1 m) wide and 3 feet (0.91 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 (centre) 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 centre 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 manoeuvre, usually the CM Pilot. The centre 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 lower equipment bay, which housed the Guidance and Navigation computer, sextant, telescope, and Inertial Measurement Unit; various communications beacons; medical stores; an audio centre; the S-band power amplifier; etc. There was also an extra rotation hand controller mounted on the bay wall, so the CM Pilot/navigator could rotate the spacecraft as needed while standing and looking through the telescope to find stars to take navigational measurements with the sextant. This bay provided a significant amount of room for the astronauts to move around in, unlike the cramped conditions which existed in the previous Mercury and Gemini spacecraft.
* The left-hand forward equipment bay, which contained four food storage compartments, the cabin heat exchanger, pressure suit connector, potable water supply, and G&N telescope eyepieces.
* The right-hand forward equipment bay, which housed two survival kit containers, a data card kit, flight data books and files, and other mission documentation.
* The left hand intermediate equipment bay, housing the oxygen surge tank, water delivery system, food supplies, the cabin pressure relief valve controls, and the ECS package.
* The right hand intermediate equipment bay, which contained the bio instrument kits, waste management system, food and sanitary supplies, and a waste storage compartment.
* The aft storage bay, behind the crew couches. This housed the 70 mm camera equipment, the astronaut’s garments, tool sets, storage bags, a fire extinguisher, CO2 absorbers, sleep restraint ropes, spacesuit maintenance kits, 16mm camera equipment, and the contingency lunar sample container.

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 (200 by 330 millimetres), used to aid in rendezvous and docking with the LM. The circular hatch window was 10 5/8 in. diameter (27 cm) and was directly over the centre 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**

* Crew: 3
* Crew cabin volume: 218 cu ft (6.2 m3) living space, pressurized 366 cu ft (10.4 m3)
* Length: 11.4 ft (3.5 m)
* Diameter: 12.8 ft (3.9 m)
* Mass: 12,250 lb (5,560 kg)
	+ Structure mass: 3,450 lb (1,560 kg)
	+ Heat shield mass: 1,870 lb (850 kg)
	+ RCS engine mass: twelve x 73.3 lb (33.2 kg)
	+ Recovery equipment mass: 540 lb (240 kg)
	+ Navigation equipment mass: 1,110 lb (500 kg)
	+ Telemetry equipment mass: 440 lb (200 kg)
	+ Electrical equipment mass: 1,500 lb (680 kg)
	+ Communications systems mass: 220 lb (100 kg)
	+ Crew couches and provisions mass: 1,200 lb (540 kg)
	+ Environmental Control System mass: 440 lb (200 kg)
	+ Misc. contingency mass: 440 lb (200 kg)

* RCS: twelve 93 lbf (410 N) thrusters, firing in pairs
* RCS propellants: MMH/N2O4
* RCS propellant mass: 270 lb (120 kg)
* Drinking water capacity: 33 lb (15 kg)
* Waste water capacity: 58 lb (26 kg)
* CO2 scrubber: lithium hydroxide
* Odor absorber: activated charcoal
* Electric system batteries: three 40 ampere-hour silver-zinc batteries; two 0.75 ampere-hour silver-zinc pyrotechnic batteries
* Parachutes: two 16 feet (4.9 m) conical ribbon drogue parachutes; three 7.2 feet (2.2 m) ringshot pilot parachutes; three 83.5 feet (25.5 m) ringsail main parachutes

## Service Module

### **Construction**

The Service Module was an unpressurized cylindrical structure, measuring 24 feet 7 inches (7.49 m) long and 12 feet 10 inches (3.91 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.

* Sector 1 (50°) was originally unused, so it was filled with ballast to maintain the SM's centre-of gravity. On the last three lunar landing (I-J class) missions, it carried the Scientific Instrument Module (SIM) which contained a package of lunar orbital sensors and a subsatellite.
* Sector 2 (70°) contained the Service Propulsion System (SPS) oxidizer sump tank, so called because it directly fed the engine and was kept continuously filled by a separate storage tank, until the latter was empty. The sump tank was a cylinder with hemispherical ends, 153.8 inches (3.91 m) high, 51 inches (1.3 m) in diameter, and contained 13,923 pounds (6,315 kg) of oxidizer.
* Sector 3 (60°) contained the SPS oxidizer storage tank, which was the same shape as the sump tank but slightly smaller at 154.47 inches (3.924 m) high and 44 inches (1.1 m) in diameter, and held 11,284 pounds (5,118 kg) of oxidizer.
* Sector 4 (50°) contained the Electrical Power System (EPS) fuel cells with their hydrogen and oxygen reactants.
* Sector 5 (70°) contained the SPS fuel sump tank. This was the same size as the oxidizer sump tank and held 8,708 pounds (3,950 kg) of fuel.
* Sector 6 (60°) contained the SPS fuel storage tank, also the same size as the oxidizer storage tank. It held 7,058 pounds (3,201 kg) of fuel.

The forward fairing measured 2 feet 10 inches (860 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 umbilical 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.

### **Service Propulsion System**

The 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. It also served as a retrorocket to perform the deorbit burn for Earth orbital Apollo flights. The engine selected was the AJ10-137, which used Aerozine 50 as fuel and nitrogen tetroxide (N2O4) as oxidizer to produce 20,500 lbf (91 kN) of thrust. The thrust level was twice what was needed to accomplish the lunar orbit rendezvous (LOR) mission mode, because the engine was originally sized to lift the CSM 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.

The propellants were pressure-fed to the engine by 39.2 cubic feet (1.11 m3) of gaseous helium at 3,600 pounds per square inch (25 MPa), carried in two 40-inch (1.0 m) diameter spherical tanks.

The exhaust nozzle engine bell 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 keep the thrust vector aligned with the spacecraft's centre of mass during SPS firings. The combustion chamber and pressurant tanks were housed in the central tunnel.

### **Reaction Control System**

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 R-4D thruster generated 100 pounds-force (440 N) of thrust, and used monomethylhydrazine (MMH) as fuel and nitrogen tetroxide (NTO) as oxidizer. Each quad assembly measured 8 by 3 feet (2.44 by 0.91 m) and had its own fuel tanks, oxidizer tanks, helium pressurant tank, and associated valves and regulators.

Each cluster of thrusters had its own independent primary fuel (MMH) tank containing 69.1 pounds (31.3 kg), secondary fuel tank containing 45.2 pounds (20.5 kg), primary oxidizer tank containing 137.0 pounds (62.1 kg), and secondary oxidizer tank containing 89.2 pounds (40.5 kg). The fuel and oxidizer tanks were pressurised by a single liquid helium tank containing 1.35 pounds (0.61 kg). Back flow was prevented by a series of check valves, and back flow and ullage requirements were resolved by containing the fuel and oxidizer in Teflon bladders which separated the propellants from the helium pressurant.

All of the elements were duplicated, resulting in four completely independent RCS clusters. Only two adjacent functioning units were needed to allow complete attitude control.

The Lunar Module used a similar four-quad arrangement of the identical thruster engines for its RCS.

### **Electrical Power System**

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, and produced drinkable water as a by-product. The cells were fed by two hemispherical-cylindrical 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.

Also starting with Apollo 14, a 400 Ah auxiliary battery was added to the SM for emergency use. Apollo 13 had drawn heavily on its entry batteries in the first hours after the explosion, and while this new battery could not power the CM for more than 5–10 hours it would buy time in the event of a temporary loss of all three fuel cells. Such an event occurred when Apollo 12 was struck twice by lightning during launch.

### **Environmental Control System**

Cabin atmosphere was maintained at 5 pounds per square inch (34 kPa) of pure oxygen from the same liquid oxygen tanks that fed the electrical power system's fuel cells. Potable water supplied by the fuel cells was stored for drinking and food preparation. A thermal control system using a mixture of water and ethylene glycol as coolant dumped waste heat from the CM cabin and electronics to outer space via 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.

### **Communications System**

Short-range communications between the CSM and Lunar Module employed two VHF scimitar antennas mounted on the SM just above the ECS radiators.

A steerable unified S-band high-gain antenna for long-range communications with Earth 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.

Four omnidirectional S-band antennas on the CM were used when the attitude of the CSM kept the high-gain antenna from being pointed at Earth. These antennas were also used between SM jettison and landing.

### **Specifications**

* Length: 24.8 ft (7.6 m)
* Diameter: 12.8 ft (3.9 m)
* Mass: 54,060 lb (24,520 kg)
	+ Structure mass: 4,200 lb (1,900 kg)
	+ Electrical equipment mass: 2,600 lb (1,200 kg)
	+ Service Propulsion (SPS) engine mass: 6,600 lb (3,000 kg)
	+ SPS engine propellants: 40,590 lb (18,410 kg)
* RCS thrust: two or four x 100 lbf (440 N)
* RCS Propellants: MMH/N2O4
* SPS engine thrust: 20,500 lbf (91,000 N)
* SPS engine propellants: (UDMH/N2H4)/N2O4
* SPS *I*SP: 314 s (3,100 N·s/kg)
* Spacecraft delta v: 9,200 ft/s (2,800 m/s)
* Electrical System: three 1.4 kW 30 V DC fuel cells

## Modifications For Saturn IB Missions

The payload capability of the Saturn IB launch vehicle 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-pound (30,300 kg) mass of the fully fuelled CSM. This was not a problem, because the spacecraft delta-v requirement of these missions was much smaller than that of the lunar mission; therefore they could be launched with less than half of the full SPS propellant load, by filling only the SPS sump tanks and leaving the storage tanks empty. 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).

The omnidirectional antennas sufficed for ground communications during the Earth orbital missions, so the high-gain S-band antenna on the SM was omitted from Apollo 1, Apollo 7, and the three Skylab flights. It was restored for the Apollo-Soyuz mission to communicate through the ATS-6 satellite in geostationary orbit, an experimental precursor to the current TDRSS system.

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, so one of the three fuel cells was deleted from these SMs.

The Command Module could be modified to carry extra astronauts as passengers by adding jump seat couches in the aft equipment bay. CM-119 was fitted with two jump seats as a Skylab Rescue vehicle, which was never used.

## Major Differences Between Block I and Block II

### **Command Module**

* The Block II used a one-piece, quick-release, outward opening hatch instead of the two-piece plug hatch used on Block I, in which the inner piece had to be unbolted and placed inside the cabin in order to enter or exit the spacecraft (a flaw that doomed the Apollo 1 crew). The Block II hatch could be opened quickly in case of an emergency. (Both hatch versions were covered with an extra, removable section of the Boost Protective Cover which surrounded the CM to protect it in case of a launch abort.)
* The Block I forward access tunnel was smaller than Block II, and intended only for emergency crew egress after splashdown in case of problems with the main hatch. It was covered with a removable plug in the nose of the forward heat shield. Block II contained a shorter forward heat shield with a flat removable hatch, beneath a docking ring and probe mechanism which captured and held the LM.
* The aluminized PET film layer, which gave the Block II heat shield a shiny mirrored appearance, was absent on Block I, exposing the light grey epoxy resin material, which on some flights was painted white.
* The Block I VHF scimitar antennas were located in two semi-circular strakes originally thought necessary to help stabilize the CM during re-entry. However, the unmanned re-entry tests proved these to be unnecessary for stability, and also aerodynamically ineffective at high simulated lunar re-entry speeds. Therefore, the strakes were removed from Block II and the antennas were moved to the Service Module.
* The Block I CM/SM umbilical connector was smaller than on Block II, located near the crew hatch instead of nearly 180 degrees away from it. The separation point was between the modules, instead of the larger hinged arm mounted on the Service Module, separating at the CM sidewall on Block II.
* The two negative pitch RCS engines located in the forward compartment were arranged vertically on Block I, and horizontally on Block II.

### **Service Module**

* On the Apollo 6 unmanned Block I flights, the SM was painted white to match the Command Module's appearance, but on Apollo 1, Apollo 4, and all the Block II spacecraft, the SM walls were left unpainted except for the EPS and ECS radiators, which were white.
* The EPS and ECS radiators were redesigned for Block II. Block I had three larger EPS radiators located on Sectors 1 and 4. The ECS radiators were located on the aft section of Sectors 2 and 5.
* The Block I fuel cells were located at the aft bulkhead in Sector 4, and their hydrogen and oxygen tanks were located in Sector 1.
* Block I had slightly longer SPS fuel and oxidizer tanks which carried more propellant than Block II.
* The Block II aft heat shield was a rectangular shape with slightly rounded corners at the propellant tank sectors. The Block I shield was the same basic shape, but bulged out slightly near the ends more like an hourglass or figure eight, to cover more of the tanks.

## In the Timeline

### **Lunar Lander (1969)**

The Apollo Command/Service Module was used in the Apollo 11 mission. Astronaut Michael Collins styed in the module whilst his collegues went to the Moon's surface. It was used to monitor the Moon from orbit as well as transporting the 3 astronauts back to Earth. Since this game did not change the events of Apollo 11 the mission went the same as it did in our timeline.

## Notes

* This page was originally taken from Wikipedia. Link: https://en.wikipedia.org/wiki/Apollo\_Command/Service\_Module