Marshall Spaceflight Center rocket designers conceived the Saturn V in 1961 and early 1960. They decided that a three-stage vehicle would best serve the immediate needs for a lunar landing mission and would serve well as a general purpose space exploration vehicle. One of the more important decisions made early in the program called for the fullest possible use of components and techniques proven in the Saturn I program. As a result, the Saturn V third stage (S- IVB) was patterned after the Saturn I second stage (S-IV). And the Saturn V instrument unit is an outgrowth of the one used on Saturn I. In these areas, maximum use of designs and facilities already available was incorporated to save time and costs.
Many other components were necessary, including altogether new first and second stages (S-IC and S- II). The F-1 and J-2 engines were already under development, although much work remained to be done. The guidance system was to be an improvement on that of the Saturn I. Saturn V, including the Apollo spacecraft, is 364 feet tall. Fully loaded, the vehicle will weigh some 6.1 million pounds. The 300,000-pound first stage is 33 feet in diameter and 138 feet long. It is powered by five F-1 engines generating 7.5 million pounds thrust. The booster will burn 203,000 gallons of RP-1 (refined kerosene) and 331,000 gallons of liquid oxygen (LOX} in 2.5 minutes. Saturn V's second stage is powered by five J-2 engines that generate a total thrust of a million pounds. The 33-foot diameter stage weighs 95,000 pounds empty and more than a million pounds loaded. It burns some 260,000 gallons of liquid hydrogen and 83,000 gallons of liquid oxygen during a typical 6- minute flight. Third stage of the vehicle is 21 feet and 8 inches in diameter and 58 feet and 7 inches long. An interstage adapter connects the larger diameter second stage to the smaller upper stage. Empty weight Off the stage is 34,000 pounds and the fueled weight is 262,000 pounds. A single J-2 engine developing up to 225,000 pounds of thrust powers the stage. Typical burn time is 2.75 minutes for the first burn and 5.2 minutes to a translunar injection.
The vehicle instrument unit sits atop the third stage. The unit, which weighs some 4,500 pounds. contains the electronic gear that controls engine ignition and cutoff, steering, and all other commands necessary for the Saturn V mission. Diameter of the instrument unit is 21 feet and 8 inches, and height is 3 feet. Directly above the instrument unit in the Apollo configuration is the Apollo spacecraft. It consists of the lunar module. the service module, the command module, and the launch escape system. Total height of the package is about 80 feet.
As the name implies, the Command and Service Module (CSM) was comprised of two distinct units: the Command Module (CM), which housed the crew, spacecraft operations systems, and re-entry equipment, and the Service Module (SM) which carried most of the consumables (oxygen, water, helium, fuel cells, and fuel) and the main propulsion system. The total length of the two modules attached was 11.0 meters with a maximum diameter of 3.9 meters. Block II CSM's were used for all the crewed Apollo missions. The Apollo 13 CSM mass of 28,881 kg was the launch mass including propellants and expendables, of this the Command Module (CM 109) had a mass of 5703 kg and the Service Module (SM 109) 23,178 kg.
Telecommunications included voice, television, data, and tracking and ranging subsystems for communications between astronauts, CM, LM, and Earth. Voice contact was provided by an S-band uplink and downlink system. Tracking was done through a unified S-band transponder. A high gain steerable S-band antenna consisting of four 79-cm diameter parabolic dishes was mounted on a folding boom at the aft end of the SM. Two VHF scimitar antennas were also mounted on the SM. There was also a VHF recovery beacon mounted in the CM. The CSM environmental control system regulated cabin atmosphere, pressure, temperature, carbon dioxide, odors, particles, and ventilation and controlled the temperature range of the electronic equipment.
The CM was a conical pressure vessel with a maximum diameter of 3.9 m at its base and a height of 3.65 m. It was made of an aluminum honeycomb sandwhich bonded between sheet aluminum alloy. The base of the CM consisted of a heat shield made of brazed stainless steel honeycomb filled with a phenolic epoxy resin as an ablative material and varied in thickness from 1.8 to 6.9 cm. At the tip of the cone was a hatch and docking assembly designed to mate with the lunar module. The CM was divided into three compartments. The forward compartment in the nose of the cone held the three 25.4 m diameter main parachutes, two 5 m drogue parachutes, and pilot mortar chutes for Earth landing. The aft compartment was situated around the base of the CM and contained propellant tanks, reaction control engines, wiring, and plumbing. The crew compartment comprised most of the volume of the CM, approximately 6.17 cubic meters of space. Three astronaut couches were lined up facing forward in the center of the compartment. A large access hatch was situated above the center couch. A short access tunnel led to the docking hatch in the CM nose. The crew compartment held the controls, displays, navigation equipment and other systems used by the astronauts. The CM had five windows: one in the access hatch, one next to each astronaut in the two outer seats, and two forward-facing rendezvous windows. Five silver/zinc-oxide batteries provided power after the CM and SM detached, three for re-entry and after landing and two for vehicle separation and parachute deployment. The CM had twelve 420 N nitrogen tetroxide/hydrazine reaction control thrusters. The CM provided the re-entry capability at the end of the mission after separation from the Service Module.
The SM was a cylinder 3.9 meters in diameter and 7.6 m long which was attached to the back of the CM. The outer skin of the SM was formed of 2.5 cm thick aluminum honeycomb panels. The interior was divided by milled aluminum radial beams into six sections around a central cylinder. At the back of the SM mounted in the central cylinder was a gimbal mounted re-startable hypergolic liquid propellant 91,000 N engine and cone shaped engine nozzle. Attitude control was provided by four identical banks of four 450 N reaction control thrusters each spaced 90 degrees apart around the forward part of the SM. The six sections of the SM held three 31-cell hydrogen oxygen fuel cells which provided 28 volts, two cryogenic oxygen and two cryogenic hydrogen tanks, four tanks for the main propulsion engine, two for fuel and two for oxidizer, and the subsystems the main propulsion unit. Two helium tanks were mounted in the central cylinder. Environmental control radiator panels were spaced around the top of the cylinder and electrical power system radiators near the bottom.
The lunar module was a two-stage vehicle designed for space operations near and on the Moon. The spacecraft mass of 15,188 kg was the mass of the LM including astronauts, expendables and 10,691 kg of propellants. The ascent and descent stages of the LM operated as a unit until staging, when the ascent stage functioned as a single spacecraft for rendezvous and docking with the command and service module (CSM). The descent stage comprised the lower part of the spacecraft and was an octagonal prism 4.2 meters across and 1.7 m thick. Four landing legs with round footpads were mounted on the sides of the descent stage and held the bottom of the stage 1.5 m above the surface. The distance between the ends of the footpads on opposite landing legs was 9.4 m. One of the legs had a small astronaut egress platform and ladder. A one meter long conical descent engine skirt protruded from the bottom of the stage. The descent stage contained the landing rocket, two tanks of aerozine 50 fuel, two tanks of nitrogen tetroxide oxidizer, water, oxygen and helium tanks and storage space for the lunar equipment and experiments, and in the case of Apollo 15, 16, and 17, the lunar rover. The descent stage was designed as a platform for launching the ascent stage from the Moon.
The ascent stage was an irregularly shaped unit approximately 2.8 m high and 4.0 by 4.3 meters in width mounted on top of the descent stage. The ascent stage housed the astronauts in a pressurized crew compartment with a volume of 6.65 cubic meters. There was an ingress-egress hatch in one side and a docking hatch for connecting to the CSM on top. Also mounted along the top were a parabolic rendezvous radar antenna, a steerable parabolic S-band antenna, and 2 in-flight VHF antennas. Two triangular windows were above and to either side of the egress hatch and four thrust chamber assemblies were mounted around the sides. At the base of the assembly was the ascent engine. The stage also contained an aerozine 50 fuel and an oxidizer tank, and helium, liquid oxygen, gaseous oxygen, and reaction control fuel tanks. There were no seats in the LM. A control console was mounted in the front of the crew compartment above the ingress-egress hatch and between the windows and two more control panels mounted on the side walls. The ascent stage was to be launched from the Moon at the end of lunar surface operations and return the astronauts to the CSM.
The descent engine was a deep-throttling ablative rocket with a maximum thrust of about 45,000 N mounted on a gimbal ring in the center of the descent stage. The ascent engine was a fixed, constant-thrust rocket with a thrust of about 15,000 N. Maneuvering was achieved via the reaction control system, which consisted of the four thrust modules, each one composed of four 450 N thrust chambers and nozzles pointing in different directions. Telemetry, TV, voice, and range communications with Earth were all via the S-band antenna. VHF was used for communications between the astronauts and the LM, and the LM and orbiting CSM. There were redundant tranceivers and equipment for both S-band and VHF. An environmental control system recycled oxygen and maintained temperature in the electronics and cabin. Power was provided by 6 silver-zinc batteries. Guidance and navigation control were provided by a radar ranging system, an inertial measurement unit consisting of gyroscopes and accelerometers, and the Apollo guidance computer.