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

On 1 September 1960 the NASA administration created the Apollo Spacecraft Project Office (ASPO) with the responsibility for the development of the Apollo spacecraft. Initially Charles W Frick and Robert O Piland headed the office to be later succeeded in November 1963 by Joseph Shea who was to oversee the initial development of the Apollo spacecraft.

One of the first contracts to be awarded by NASA was to the Massachusetts Institute of Technology (MIT) for the Guidance and Navigation of the spacecraft. The task of developing the Apollo navigational systems fell to the Instrumentation Laboratory of MIT and its director Charles Stark Draper, who were asked in November 1960 to conduct a feasibility study which was followed by a letter contract for both hardware and software on 9 August, 1961. Their experience stemmed from being the prime contractor for production of a guidance computer for the 'Polaris' ICBM. The Apollo Guidance Computer (AGC) was designed and developed by MIT and was to be manufactured by the Raytheon Corporation of Massachusetts.

The prime contract for the Apollo spacecraft itself, went to North American Aviation who were also to be responsible for the integration of the lunar module and launch escape systems within the overall concept. They were awarded the contract over four other competitors on 28 November 1961 and subsequently agreed an initial contract value of 934.4 million dollars with NASA in August 1963. NAA was acquired by the Rockwell Manufacturing Company and became the North American Rockwell Corporation in September 1967. North American's Space and Information Systems Division was headed by NAA vice president Harrison A 'Stormy' Storms and included in their Apollo design team John Paup as program manager, Norman J Ryker Jnr as chief designer and Charles H Feltz who already had links with NASA from his work with the X-15 rocket plane production.

Grumman Aviation, a company with a long and proud tradition of manufacturing aircraft for the US Navy who had conducted feasibility studies on the LOR mission mode, won the contract for the lunar module on 14 January, 1963, after just losing to North American for the CSM contract a year earlier. The contract was formally signed in March 1963 for a projected cost of 387.9 million dollars.

The Apollo spacecraft was a modular assembly comprising of two main spacecraft components, a Command and Service Module (CSM) and a Lunar Module (LM). The CSM housed and sustained the three-man crew during the major part of the lunar journey, while the LM was used for the descent to a landing by two of the crew, and their return to the CSM, which remained in lunar orbit during the landing. Both the CSM and the LM each consisted of two modules with individual functions that permitted a considerable weight saving advantage to be gained by being able to discard those modules that became redundant after use.

The CSM consisted of the Command Module (CM) and the Service Module (SM), which were mated together during manufacture and remained attached for all but the last few hours of the mission. The CM housed the three man crew from where the mission was conducted. The SM provided electrical power, water, communication and propulsion for the command and was jettisoned just prior to re-entry on return to earth. The LM comprised a descent stage and an ascent stage, each with their own engines. The descent stage was to be used for the descent from lunar orbit to a landing and as a stable take-off platform for the upper, ascent stage. The ascent stage, perched atop the descent stage, housed the two man crew during the landing and returned them to the CSM.

Both parts of the Apollo spacecraft sat atop the S-IVB third stage. The LM was housed immediately above the S-IVB third stage in a Lunar Module Adapter (LMA), a conical frame of four fairing panels which also protected the LM from aerodynamic forces during the launch phase. The CSM was perched above the lunar module atop the adapter. This configuration of the CSM above, and separate from, the LM was used during launch and while in earth parking orbit. This provided the facility that in the event of an emergency abort of the mission, the command module with its crew, being uppermost on the stack, could be pulled clear of the remainder of the spacecraft without hindrance, by a rocket mounted on a tower above the command module.

This configuration also meant that the CSM and the LM had to be mated together after launch to carry out the mission. Connection of the two spacecraft was achieved by the Transposition and Docking Manoeuvre (TDM) carried out by the CM pilot after the TLI burn using a drogue and probe mechanism. The Apollo Docking Mechanism (ADM) consisted of a probe housed at the forward end of the command module's apex, inside a docking tunnel, that engaged with a dish shaped drogue in the upper docking hatchway of the lunar module. On insertion of the probe into the drogue, three capture latches engaged with a hole in the drogue's apex to form a 'soft dock'. Firing a helium gas charge operated a retraction mechanism of the probe which pulled the two craft together so that twelve latches in the command module's docking ring, which surrounded the probe, could engage with a corresponding ring in the lunar module to form a 'hard dock'. Crew transfer between the two spacecraft was possible through the docking tunnel after removal of the CM's forward hatch, the probe and drogue and the LM's upper hatch.

Although the Saturn V's payload capability was considerable, it was not unlimited. Weight considerations and reliability for the Apollo spacecraft were of paramount importance. These considerations affected the design of the spacecraft and forced new procedures and design techniques. One of the main weight savings was obtained by the use of a pure oxygen atmosphere for the crew in the spacecraft. Although earthly atmospheric pressure is 14.7 pounds per square inch and consists approximately of two-thirds nitrogen and one-third oxygen, transposing that internal pressure into space would require a strong and heavy construction to contain it in an external vacuum. Using a pure oxygen atmosphere required only an internal pressure of just under 5 psi for the crew's needs, which in turn only required a significantly lighter construction to retain it.

The decision to utilise a pure oxygen atmosphere was not without significant consequences for the crews. It required them to pre-breath pure oxygen for some hours before launch. This was to remove the nitrogen content from their circulatory systems in order to prevent nitrogen narcosis, (bends) when the spacecraft's internal atmospheric pressure dropped as it gained altitude. It was also to have catastrophic effects for the crew of Apollo 1 during a countdown test.

Almost all equipment within the spacecraft had a back-up system or a built-in redundancy factor whereby failure of one system could be over-ridden or compensated for, by the use of another. However, design and weight considerations of certain critical equipment meant they could not be duplicated and consequently required absolute reliability. In particular, the main engine of the Service Propulsion System (SPS) in the service module which would be required for up to eight restarts and be used to place the spacecraft into a lunar orbit and more importantly, get it out again. This also applied to the lunar module's single use, ascent engine, used for return from the lunar surface. Failure of either of these two engines would result in a crew trapped in orbit or on the moon's surface.

Working toward the goal of simplicity and thereby reliability, all three of the main engines in the Apollo spacecraft used hypergolic (self igniting when combined) fuels in a heady, 50-50 mixture of unsymmetrical dimethyl hydrazine and hydrazine propellant, with a nitrogen tetroxide oxidiser. Fuel systems were pressurised by helium to provide fuel flow without the need for complex pumps and ignition systems while pressurisation initiation was provided by pyrotechnic 'squibs' used to open single use valves.

Overheating or freezing of the spacecraft during flight was regulated by the adoption of Passive Thermal Control (PTC), otherwise known as 'barbecue mode'. The spacecraft was made to roll around its longitudinal axis at a rate of 0.3 degrees per second. This exposed all the external surface area of both craft during a period of about 20 minutes preventing any one part of the spacecraft being exposed to prolonged heating in sunlight, or freezing in shadow. The use of a reflective Mylar covering on the spacecraft's outer surface helped to prevent overheating from exposure to the sun's radiation.

Thermal control of internal electronic equipment which produces heat in the normal course of its operation was maintained by mounting electronic equipment on heat sink rails, that were in turn cooled by a water/glycol mixture that circulated through the rails. The glycol coolant circulated in a primary feed loop carrying heat away from the heatsink to radiators incorporated in the spacecraft's outer skin. A secondary stage of cooling was obtained by the use of sublimators where water from an open, secondary feed, removed heat from the primary loop and was evaporated into space.

Launch Escape Tower (LET)

Topping off the whole of the Apollo stack was a Launch Escape Tower (LET) which incorporated a Boost Protective Cover (BPC). The cover fitted closely over the conical shaped leading face of the CM insulating it from the 1200 degree heat that would be generated by air friction during the initial flight through the lower part of the earth's atmosphere. Developed by North American Aviation, the function of the LET as part of the Abort Escape System (AES), was to provide a means of escape for the crew in an emergency from five minutes before lift off, until three minutes into the launch. Three minutes after lift off, just after first stage separation and when clear of the earth's atmosphere where parachutes would be of no further use, the LET would be fired to jettison the boost protective cover clear of the command module.

The LET contained three solid fuel rocket motors mounted on a lattice tower framework. A 147,000 pound thrust rocket, which could be triggered either automatically by the launch computer, or manually by the spacecraft commander, could pull the command module and its occupants free of the stack and return them to earth by parachute. The tower also incorporated a separate, 31,500 pound thrust jettison motor to pull the tower and boost protective cover clear of the spacecraft and a small 2,400 pound motor to provide pitch control during an abort sequence. These motors were all solid fuel rockets and were manufactured by the Lockheed Propulsion Company and the Thiokol Chemical Company. The LET also provided atmospheric pressure sensors at its forward tip to supply data to the IU. During countdown to launch the sensor apertures were covered by a Q-ball on the uppermost arm of the umbilical tower which was retracted few minutes before launch.

The abort escape system underwent a number of tests at the US Army's White Sands missile testing range in the New Mexico desert. Simulated launchpad escapes from a static testbed and in flight to heights of 180,000 ft, using a booster rocket 'Little Joe II', tests were carried out to ensure its effectiveness. The system functioned as predicted but was never tested with live occupants.

In June 1968, manned tests of the CSM and the LM were carried out in large vacuum chambers at the Space Environment Simulations Laboratory (SESL) at Houston. Two chambers had been prepared big enough to individually house the spacecraft and simulate the heat, cold and vacuum conditions that would be experienced in space. On 16 June, 1968, astronauts Joe Kerwin, Vance Brand and Joe Engle entered command module spacecraft 2TV-1 (Thermal Vacuum Test) to spend 177 hours 'flying' the test vehicle inside the vacuum chamber to prove the viability of the re-designed spacecraft. Their mission patch included the depiction of a road-runner bird and the motto 'Arrogans Avis Cauda Gravis' (The bird with the heavy tail). A second manned test of the lunar module in the vacuum chamber was carried out by astronaut James Irwin and Grumman test pilot Gerry Gibbons who spent 48 hours in Lunar Test Article LTA-8.