Project Apollo: The Beginnings | Mission Planning | Landing Site Selection | Earthbound Support Systems
Astronaut Selection and Training | The Saturn/Apollo Stack | 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

After much consideration and internal argument, NASA had by 1962 adopted the Lunar Orbit Rendezvous (LOR) mode as the most likely means to meet the 'end of decade' deadline set by President Kennedy's challenge to '...land a man on the moon and return him safely to earth'. This technique dictated the design of the craft to be used for the landing attempt and the specification of the launch vehicle. LOR required the placing a two part spacecraft into orbit around the moon from where one half, a lunar module, would descend to a landing leaving the other half, a command and service module, in orbit. On completion of the lunar excursion the lunar module would take off and return to the orbiting command ship.

While undergoing assembly and readying for launching, the complete Saturn/Apollo moonship was referred to by its builders as 'the stack'. It was comprised of the two part Apollo spacecraft at the top of a three stage Saturn V launch vehicle. On the launchpad the complete Saturn/Apollo stack stood 363 feet high and when fully fuelled weighed over 3000 tons. The individual component assemblies that comprised the stack were each the product of separate manufacturers located throughout the United States and were brought together for assembly in the Vehicle Assembly Building (VAB), at the Kennedy Space Centre (KSC) launch site in Florida, USA. The component manufactures diverse locations ranged throughout the United States from New York State to Alabama and California.

The Saturn V Launch Vehicle

The Saturn V launch vehicle used to get the Apollo spacecraft into earth orbit and on its way to the moon was assembled from a S-IC first stage booster, a S-II second stage and a S-IVB third stage. Throughout the Apollo program thirteen Saturn Vs were launched without loss. Each vehicle took five months from the time the first components were delivered to the Kennedy Space Centre (KSC) to assemble, prepare, test and launch. In 1969, at the height of the Apollo program, three vehicles were in various stages of assembly or ready for launch, but only once was the second launch pad at KSC used. This was for the launch of Apollo 10 from pad 39B, all other Saturn V's were launched from pad 39A.

The Saturn launch vehicle design and development was the responsibility of NASA Marshall Space Flight Centre (MSFC), Huntsville, Alabama under its Centre Director, Dr Wernher von Braun. Each of the vehicle stages, instrument unit and engine programs had individual management teams and construction took place at the individual manufacturers plants.

Saturn S-1C First Stage

The first stage Saturn S-IC booster was constructed by the Boeing Co at its Michaud facility near New Orleans. It was 33ft in diameter and 138ft tall and was designed to initially lift the complete spacecraft through the lower part of the earth's atmosphere to an altitude of over 40 miles. It was the most powerful engine ever built and successfully launched. The S1-C and its fuel load made up two thirds of the stack's overall launch weight. The S-IC was capable of lifting the total launch weight and accelerating it through the speed of sound in a near vertical climb to reach a speed of nine times the speed of sound at engine cut off just three minutes after lifting off. With weight considerations of crucial importance the construction of the stage, in common with the remainder of the other two upper stages, was principally from an aluminium alloy developed by the Alcan Co. Constructed from a composite ring frame and stringer assembly, no part of the frame material used were thicker than 0.25 inches.

Two main propellant tanks formed part of the thrust and weight bearing load structure and were connected by an inter-tank section. The upper tank contained 345,000 gallons of liquid oxygen (LOX) and was initially pressurised by helium on the launchpad, but transferred to gaseous oxygen (GOX) after lift-off which was supplied by a heat exchanger drawing LOX from the fuel supply. The lower tank containing 203,000 gallons of a refined kerosene mixture (RP-1) and was pressurised with helium from a ground supply on the pad and by four bottles housed inside the LOX tank during flight. The construction of the tanks was so light that they had to be kept pressurised to prevent them from buckling under their own weight.

At the base of the stage a thrust frame mounted the five Rocketdyne F1 engines required to lift the spacecraft. The S1-C's engines were developed and built by Rocketdyne, a subsidiary of North American Rockwell Corporation. Designated the F1, each engine weighed ten tons and stood 18 feet tall with an exhaust bell exit diameter of 14 feet. Turbopumps supplied the cluster of five engines, which consumed between them over fifteen tons of fuel and oxidiser each second to develop a total thrust of seven and a half million pounds. The central F1 engine's alignment was fixed but the surrounding four engines were mounted to the thrust frame and could be pivoted up to six degrees to provide directional thrust and steering.

A stainless steel, honeycombed lower heatshield situated between the engine thrust frame and the lower propellant tank safeguarded heat critical components from the engines. Four titanium/aluminium shrouds incorporating fixed stabilising fins protected the outer engine bells from aerodynamic forces and housed the engine gimballing linkages. Also contained within the shrouds, grouped in pairs, were eight, solid fuel retro-rockets, each producing 87,900 pound thrust for a duration of 0.6 seconds when fired one second after the stage was jettisoned. These would slow the S-1C to provide separation from the second stage as its engines ignited.

Interstage Ring

An annular interstage ring separated the first and second stages and aerodynamically faired in the gap between them while providing clearance for the second stage engine bellmouths. The ring also housed four 219,000 pound thrust, solid fuel rockets, which were fired just after the first stage was jettisoned when acceleration had momentarily ceased. This provided 'ullage' of the fuel in the second stage's tanks, forcing the fuel to the bottom of the tanks ready for ignition of the second stage engines. The interstage ring was jettisoned 30 seconds after the S-1C and was referred to as a 'two plane separation'.

Saturn S-II Second Stage

The second stage Saturn S-II, built by North American Rockwell Corporation at Seal Beach California, maintained the overall diameter of the first stage at 33 feet but was 81 feet in length. As the S-IC exhausted its fuel and was jettisoned at an altitude of 220,000 feet the S-II took over to accelerate the remainder of the spacecraft, now weighing only one third of its original weight, through the upper atmosphere to a near orbital altitude of 610,000 feet.

Using five Rocketdyne J2 engines, burning more efficient cryogenic fuels, the engines developed over one million pounds of thrust between them. 260,000 gallons of liquid hydrogen and 83,000 gallons of liquid oxygen were housed in tanks separated by a common bulkhead. The mounting of the engines was similar to that of the first stage which allowed gimballing of the outer four engines for steering while the central engine's alignment remained fixed.

The S-II also housed four 219,000 pound thrust solid fuel ullage rockets to settle fuel in the second stage tanks. The difference in diameters between the second and third stages was accommodated by an interstage ring which tapered the divergent diameters and provided a step in the spacecraft's outline. Four further retro-rockets housed in the top of the interstage ring provided clean separation of the second stage from the third in a similar manner to the S-1C. This interstage ring was jettisoned still attached to the second stage.

During its development the design of the S-II proved to be the most troublesome of the three stages. Initially the fuel tanks proved liable to cracking and were redesigned. It also had a tendency to resonate on its longitudinal axis causing repeated oscillation of the framework which give rise to the phenomenon's name, 'pogo'. Modifications to the J2 engine's fuel supply system and bringing the centre engine cut-off to a minute and a half before staging, reduced the effect bringing it within acceptable limits, although the effect was never fully eradicated.

S-IVB Third Stage

The third stage Saturn S-IVB was built by McDonnell Douglas Astronautics Co at its Huntingdon Beach, California plant. It was 22 feet in diameter and 58 feet in length and powered by a single, restartable Rocketdyne J2 engine. Burning liquid hydrogen and liquid oxygen to produce almost 250,000 pounds of thrust, the S-IVB would increase the spacecraft's launch speed to 17,400 miles per hour to provide the final push to orbital speed and height. The S-IVB's second function was to carry out the Trans Lunar Injection (TLI) manoeuvre. When safely checked out in earth orbit and ready to commit to a lunar mission, it had the ability to restart its engine and accelerate the Apollo spacecraft to an escape velocity of 24,200 mph that would take it out of earth orbit and onto a path to the moon.

Pressurisation of the LOX tanks came from eight spherical helium tanks inside the liquid hydrogen tank, while ullage was provided by two further solid fuel rockets housed in external pods. Attitude control was provided by the Auxiliary Propulsion System (APS), eight liquid fuel thrusters in two propulsion system modules. The S-IVB was also used as the second stage for all earth orbital missions which used the smaller, less powerful, Saturn S-IB first stage booster.

Instrument Unit (IU)

Attached to the top of the S-IVB housed in an three foot high annular ring, mounted at the top of the S-IVB, was the guidance and control system Instrument Unit (IU). Manufactured by IBM Federal Systems Division, it contained the on-board Launch Vehicle Digital Computer (LVDC), which controlled the Saturn's flight guidance and navigation systems from lift off to earth orbit and throughout the Trans Lunar Injection (TLI) engine burn. Its functions included power supply and cooling to the Saturn's electrical and electronic systems, guidance of the launch vehicle and monitoring performance and fault diagnosis. Its communication system also returned data to the ground based mission control computers to allow ground based monitoring of its condition and performance.

Initially the IU's orientation was set with a visual reference point on the ground, and transferred to internal battery power 17 seconds before lift off, then released from ground control with five seconds to go. Monitoring the orientation of the Saturn was accomplished by a ST-124 inertial platform, manufactured by the Bendix Corporation, that measured its change of attitude through the three axes of roll, pitch and yaw while changes in the speed of the craft were measured by accelerometers. Data from the platform and other sensors was used by the LDVC to compute course corrections and adjust the engine's alignment to optimise the crafts trajectory. The IU also incorporated an Emergency Detection System (EDS) which in automatic mode could initiate the Abort Escape System (AES) abort sequence to terminate the flight if necessary. All three stages had shaped explosive charges attached to the main fuel tanks, controlled by the IU, which could be used in the event of an abort to rupture the tanks to disperse the remaining fuel. As a back up to total failure of the LVDC, the command module guidance computer which monitored the LVDC, could take over flight guidance during the launch.

Saturn 1B

The early Apollo development flights including launches into earth orbit of manned and unmanned test vehicles. It required the power of the Saturn V to launch a CSM and a LM together but when the mission required only an unmanned or manned test flight into earth orbit the smaller Saturn 1B was used as a first stage with the Saturn S-IVb as a second stage. The Saturn 1B was built by the Chrysler Corporation at the Michaud assembly facility and used a cluster of eight Rocketdyne H1 liquid fuelled engines with a combined thrust of 1.6 million pounds. Apollo missions using the S-1B first stage booster were launched from pad LC-34 on the Cape Canaveral launch site

Apollo launches using the S-1B were:

• AS-201 Launch vehicle development. Sub-orbital
• AS-202 Launch vehicle development. Command module heat shield test
• AS-203 Launch vehicle development. Fluid dynamics test
• Apollo 5 Unmanned test of the Lunar Module
• Apollo 7 Manned test of the Command and Service Module
• Skylab 2 Launched first crew to Skylab Space Station
• Skylab 3 Launched second crew to Skylab Space Station
• Skylab 4 Launched third crew to Skylab Space Station
• Apollo-Soyuz Launched crew for rendezvous with Soviet Soyuz Spacecraft.

The Apollo Spacecraft

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 were each comprised 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, comprised 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 of 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 decompression illness (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.

Guidance and Navigation (G and N)

While many of the aspects of the construction of the Apollo and Saturn hardware was based on existing technology, one aspect of the project, the spacecraft's guidance and navigation required extensions into new disciplines. To place a spacecraft into orbit and track it around the earth was one thing, but to send it off beyond the gravitational influence of the earth and have it arrive at a precisely defined point just 60 miles in front of another planetary body, at a precise speed and time would require some original thinking.

It was considered desirable to provide the crew with the facility to input their own navigational data in the event of an emergency situation or loss of contact with the ground. Some of the manoeuvres necessary for a manned landing by the LM and initiation of the Trans Earth Injection (TEI) manoeuvre would have to be initiated by the crew out of sight of the earth. This would require a degree of autonomous control and the need for an on-board computer to be capable of controlling the navigation, flight guidance and control computations without earthbound assistance.

As both the CSM and the LM were at times required to operate independently of each other during the mission, each craft was equipped with its own independent means of navigation. Although the individual computers were practically identical, the command module's computer would handle guidance between the earth and the moon and in lunar orbit, while the lunar module's provided guidance during the lunar landing, ascent and rendezvous. The software code differed for the separate purposes and that used in the command module was named 'Colossus' and a second program, 'Luminary' was used in the lunar module.

Command Module (GandN)

The G and N system of the Apollo spacecraft comprised of three main subsystems, computer, inertial and optical. The Apollo Guidance Computer (AGC), kept a running log on the spacecraft's velocity and position as it changed due to gravitational effects of the earth and moon, matching it with a pre-planned flight trajectory. The AGC received data from its own Inertial Measurement Unit (IMU), measuring rotation of the spacecraft around the three axes of pitch, roll and yaw and from accelerometers which measured changes in the spacecraft's velocity during engine firings.

Inevitably, due to precession in the IMU's gyroscopes, alignment tended to drift off and required periodic resetting. This was achieved through the spacecraft's optical system (OPS), by taking two star sightings through its sextant and comparing readings with a known list of 37 stars who's positions were stored in the computer's memory. This fix, when input to the AGC would establish the spacecraft's attitude.

The computations and control commands during a lunar flight were to be carried out by ground based equipment at MCC and uplinked to the spacecraft. By measuring the time taken to transmit and return a ranging signal and its Doppler shift, the spacecraft's range and speed could be determined with a high degree of accuracy. This data was included in a mathematical computation, a 'state vector', which was transmitted to the spacecraft and input to its guidance computer which enabled the establishment of the spacecraft's position and speed.

Crew input into the AGC was through the Display and Keyboard system (DSKY) a 21 digit display panel and 19 pressbutton keyboard, referred to as the 'disky'. Selection of programmed manoeuvres and their initiation was controlled by the crew's input of two figure numerical commands, which called up the required program from the computer and the action required.

CM Flight Control

The G and N system was directly linked to the spacecraft's Stabilisation and Control System (SCS). Computations made by the AGC provided a 'thrust vector' to align the main engine for major velocity and direction changes and to control the spacecraft's rotation about its axes. The SCS also permitted manual control of spacecraft attitude through two translation (hand) controllers which allowed the CM Pilot to fire external thrusters to manoeuvre the spacecraft.

Lunar Module Primary Guidance Navigation and Control (PGNCS)

Once separated from the CSM, the LM had to navigate its own way to the lunar surface and on return, manoeuvre to a rendezvous with the waiting CSM. For this purpose it carried two guidance systems to manage these complex manoeuvres. The Primary Guidance Navigation and Control System (PGNCS), referred to as 'pings', utilised its own Lunar Module Guidance Computer (LGC), which was essentially a duplication of the AGC computer used in the command module. Manual input to the computer was through a DSKY, located at the commander's station.

To manoeuvre the spacecraft, the PGNCS computer automatically controlled thrusters mounted externally on the ascent stage and the descent engine's thrust and alignment to maintain the desired flight path, through a series of programs entered by the crew through the DSKY. The LGC received data from its own inertial guidance platform that could be updated through its optical star sighting system. During landing the computer also received data from a landing radar sub-system, which became operative from approximately 40,000 feet above the lunar surface, providing height and rate of descent of the LM. During the LM's ascent, a separate rendezvous radar updated the LGC after locating the CSM at ranges of up to 350 miles and provided range and rate of closure with its target.

The landing radar, rendezvous radar and stabilisation and control system were supplied by Aerospace Communications and Control, a division of the Radio Corporation of America (RCA). RCA designed and manufactured the rendezvous radar and bought in the landing system from Ryan Aeronautical Co who had previously developed the landing radar for the Surveyor soft landing lunar probes.

The Abort Guidance System (AGS)

In the lunar module the PGNCS was backed up by a separate, independent computer, the Abort Guidance System (AGS) referred to as 'aggs'. The AGS incorporated its own strapped-down inertial system and with data from the landing or rendezvous radar providing a running check on the PGNCS function. In the event of a PGNCS failure it provided guidance and control for the ascent stage back to rendezvous with the CSM, or an orbit from which it could be rescued. Input to the AGS was through its own DSKY at the lunar module pilot's position. It was never found necessary to use the AGS on any of the lunar missions except under test conditions.

The Command and Service Modules CSM

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.

The Lunar Module (LM)

An old saying in aviation circles says, 'If it looks right, it will probably fly right', but based purely on functionality, a more unlikely looking flying machine would be difficult to imagine. Designed and manufactured by Grumman Aerospace Corporation at Bethpage, New York State, it consisted of two stages, the descent, and the ascent stages.

The Lunar Module (LM) was originally known as the Lunar Excursion Module until the 'excursion' part of the name was deemed too frivolous and dropped. In its early life it was referred to occasionally by NASA as the 'bug', by its manufacturers as the LM, but nearly always by the astronauts as the 'lem'. It was originally designed to operate independently for only 48 hours and its sole function, once separated from the CSM, was to descend from lunar orbit with its two man crew to a landing, sustain them during their stay and return them to orbit to rejoin the CSM.

From its inception it was subject to weight and development problems that made it the pacing element of the Apollo program. Grumman's original concept by the design team headed by Thomas J Kelly was subjected to changes of design during its evolution. The first conception of the LM, on which the contract was won, sported five fixed landing legs, two docking hatches and the crew seated behind panoramic glass windows similar to those of a helicopter. Power to the LM's electrical systems was to be derived from power cells similar to those in the CSM and crew egress to the lunar surface was via a knotted rope hanging down the side of the LM.

One of the first major changes reduced the number of landing legs to four. The original concept of a circular configuration was not sufficiently robust to withstand the loads imposed during launch on the adapter housing of which the LM descent stage formed a structural part. The descent stage was modified to a cruciform configuration and the landing legs naturally fell to one at each corner. This change dictated the use of four larger propellant tanks from the originally envisaged six, which gave a weight saving and a reduction in the amount of plumbing. NASA's considerations with regard to the lunar soil bearing strength demanded larger landing footpads and the new design would not fit within the confines of adapter housing, consequently this required the design to be modified to provide legs that could be extended from a folded position in flight.

It was also soon realised that the internal living space of the ascent stage was not going to be sufficient to include NASA's increasing inventory of necessary equipment. EVA suits with backpacks, helmets and sample return containers could not all be accommodated within the projected LM's living space. Realisation that the crew did not need to sit down during the landing resulted in the deletion of seats giving a significant increase of the internal space. This also had an added advantage that the crew would be moved closer to the windows allowing a reduction in the size of the window for the same field of vision. This reduction of the area of glass resulted in a significant weight saving and the removal of the potential hazard of a structurally weak material. All that was necessary for the crew's security was to provide a spring loaded harness to waist belts and foot cleats. Further fold-down armrests and hand holds on the control panel allowed the crew to brace themselves against the harness while operating the spacecraft's attitude hand controller.

Experiments with crew egress established that exiting the front docking hatch in a bulky spacesuit and large backpack to lower himself to the lunar surface by a knotted rope, was impractical even in one sixth gravity. This resulted in the deletion of the forward hatch docking mechanism and reshaping it to an oblong configuration, resulting in a small but adequate aperture. The addition of a ladder on the front landing leg, accessed from the small 'porch' outside the hatch, improved access to and from the living quarters of the ascent stage.

As the design progressed, so the craft's weight increased. The original concept weight was 22,000 pounds which was increased in the contract to 25,000 pounds. In January 1964 the design weight was increased to 29,500 pounds and again in November 1964 to 32,000 pounds as changes and additions were forced into the design. By the end of the year estimates projected that even that would be exceeded as the craft's weight was still escalating. In July 1965 Grumman instigated a weight saving campaign, 'Operation Scrape' which resulted in a marginal saving . A further 'Super Weight Improvement Program'( SWIP) at NASA's insistence, subjected the LM to a major weight reduction program involving not only Grumman but also sub-contractors for every part supplied for use in the LM. By the end of 1966, SWIP had succeeded in shaving 2,500 pounds from the lander's weight to keep it just within its design limit and provision for further development.

Several design changes were necessary because of difficulties with the development of the new technology. In the original concept electrical power in the LM was to be produced by power cells similar to those used in the CSM. Difficulties with the LM's power cells resulted in a change to silver-zinc batteries. Although this system was simpler and removed the need for oxygen and hydrogen tanks and their associated plumbing, the net result was an increase in the LM's weight. Neither did the change to batteries immediately bring the hoped for increase in reliability. During testing the battery's power output was found to be erratic and unpredictable. It was not until an investigation by Grumman with the battery manufactures revealed production weaknesses that the problem was solved.

Major problems began to develop in 1967 and 1968 at a late stage in the LM's development. Chemical milling , a process used extensively in the production of alloy parts to remove excess weight by treatment with acids, increased the possibility of the part cracking under stress. It was found that unless the parts were a near perfect fit the stresses induced when fitted up, could cause the part to crack, even months after assembly. All accessible parts on the LM's then under construction for manned flight were examined and some of the major components were replaced while new production procedures were introduced to minimise the problem.

Further fuel supply problems plagued Grumman throughout development. Due to the highly corrosive nature of the engine propellants being used for both of the LM's engines and the RCS, joints within the fuel supply systems were prone to leakage. With the fuels being hypergolic any leakage and mixing of the propellants was highly undesirable within the LM's structure. Redesign of seals for the propellant tank connections and introduction of welded joints where-ever possible eventually rectified the problem. Production changes were also introduced requiring each individual joint to be x-rayed and certified.

The combined result of these and many other problems resulted in the lunar module significantly lagging behind the rest of the Apollo program. As with North American, relations between NASA and Grumman were at times strained, not least on the late delivery of the first flying lunar module LM-1. Despite a favourable acceptance review a week earlier, LM-1 failed NASA's fuel leakage tests when received at KSC and prompted Rocco Petroni, NASA's Director of Launch Operations, to comment to Grumman design staff '...it's a piece of junk... garbage'. Further wiring and fuel leakage faults found in LM-3 was a significant factor in the decision to extend the Christmas 1968 Apollo 8 mission from a high altitude test to a circumlunar flight of the CSM.

Descent Stage

The descent stage, roughly octagonal in shape measuring 14 feet 1 inch corner to corner or 31 feet with its legs fully extended in flight. It was of cruciform construction housing the main descent engine in its central bay with four fuel tanks at opposing sides. The landing gear which consisted of four extendible legs were attached to the main frame at 90 degrees to each other. The legs incorporated an internal collapsible honeycomb that deformed on contact with the lunar surface to provide shock absorption for a descent rate of up to ten feet per second (fps) vertically, or seven fps when combined with a four fps sideways drift. Each of the legs ended in a 37 inch diameter circular pad that was to provide a firm footing on the expected loose, dusty surface. Extending under three of the pads were 68 inch long probes that tripped a blue warning light on the instrument panel when making contact with the lunar surface, giving the pilot time to shut down the engine as the craft settled onto the surface. The probes collapsed under the pads on contact.

The outer framework of the descent stage was covered in a single alloy sheet and multiple layers of Mylar foil with a reflective coating that gave the covering an overall gold colour. The use of Mylar was a result of the weight saving campaign, as it was light and its reflective qualities gave excellent thermal protection to internal heat sensitive mechanisms, while its multiple layers provided protection against micro meteorites. Once on the moons surface the descent stage also formed a stable launchpad from which the upper half of the LM, the ascent stage, could take off.

The specification for the Descent Propulsion System (DPS) included a variable thrust requirement to allow control of the engine during the descent manoeuvre where maximum thrust would be required to slow the spacecraft to allow it to descend from orbit and decreasing thrust as the craft pitched over to allow a near vertical descent in the final phase before touchdown. Throttled rocket engines had not been previously used in spaceflight and Grumman commissioned the simultaneous development of two separate engines by different manufacturers, using different methods of throttling. One engine, developed by Rocketdyne used inert helium gas injection into the fuel supply to limit thrust. The other, by Space Technology Laboratories (STL), used mechanical throttling by a restrictive valve system, coupled with a variable fuel injector head to limit fuel flow to the combustion chamber.

Both engines developed equally well and in January 1965 STL's engine was chosen for the lunar module. It was a throttled, gimballed rocket motor that produced from 1,050 to 9,870 pounds of thrust. It was gimballed to provide directional thrust up to six degrees for manoeuvring and to compensate for variations in the spacecraft's centre of gravity as the fuel load was consumed. Thrust and alignment were controlled by the PGNCS.

Also situated within the descent stage, right hand corner quadrant, and adjacent to the access ladder, was the Modularised Equipment Stowage Assembly (MESA) bay, in which was stowed equipment for the lunar stay. Tools, replaceable consumables such as the air filters used in the ascent stage ECS and the surface television camera were accessed by a drop-down flap which also doubled as a work table to the MESA. On the opposite quadrant, to the rear of the craft, provision was made for housing the Apollo Lunar Surface Experimental Package (ALSEP). This was a packages of instruments and experiments to be left on the moon's surface and was accessed under a peel-away Mylar covering. On later missions, a collapsible, electrically driven car, the Lunar Rover, dubbed the 'moon buggy', would also be housed on the descent stage framework.

Ascent Stage

The ascent stage carried the two man crew in a pressurised compartment with their life support systems, control, navigation and communication systems and EVA suits. It was powered by its own ascent engine and fuel for the return trip to lunar orbit was contained in two spherical tanks on either side of the crew compartment. The weight of the fuel contained within the tanks meant that the oxidiser tank had to be mounted further outboard to maintain the craft's weight distribution. When viewed face-on it gave the spacecraft an unsymmetrical look and was likened to '... a Hamster with mumps on one cheek'.

Flight guidance manoeuvring of the ascent stage during the ascent, was solely through the Reaction Control System (RCS) which mounted four quads of thrusters on outriggers at the front and rear of the stage. Control of the RCS was through the PGNCS computer and propellant for the system was stored in spherical tanks on the exterior of the crew compartment. Manual control of the RCS for final docking could be taken by either crewmember who were each provided with translation hand controllers. Provision was made for the RCS propellant tanks to be topped up from the descent engine fuel tanks.

Crew Compartment

The crew compartment was cylindrical in section in a welded and riveted construction, 92 inches in diameter and 42 inches deep, giving a habitable volume of 160 cubic feet, just sufficient for the two crewmembers to stand side by side. Due to the weight saving programs the compartment skin was reduced to a thickness of 0.012 inches, the equivalent of approximately three layers of kitchen foil. The crew were restrained in a standing position by spring loaded straps to the side of the compartment and its floor.

Access to the compartment from the CM was from the overhead hatch to the docking tunnel. A forward facing hatch placed centrally under the control panel allowed egress from the crew compartment to the external porch and ladder on the descent stage. The hatch incorporated a dump valve to allow the internal atmosphere to be vented to space for the EVA and was hinged on the right to open inwards. Due to the limited interior space this made it necessary for the commander on the left to exit first as it was impossible for the lunar module pilot (LMP) to get through the hatch while the commander's position was occupied. It also meant that the commander had to be last in.

At the forward end of the crew compartment two downward tilted, triangular observation windows allowed the crew a forward view. Each window was triple paned and optically coated. The left hand, commanders window, was etched with graduations that would be used to determine the landing spot. Two graduated scales could be 'eyeballed' into alignment and using reference data called out by the LM Pilot, the commander could see the exact point on the ground to which the guidance computer was taking them. The control panel and DSKY was centralised between the two windows with the optical alignment telescope above the instrument panel at eye height. A further small rectangular docking window was let into the roof of the compartment to allow the commander a view of the approaching CSM during docking manoeuvres.

A raised cover over the ascent engine took up space to the centre-rear of the cabin between the crewmembers while the rear of the compartment was taken up by the Environmental Control System, storage space for EVA suits, backpacks, helmets, food and equipment. Rest periods while on the lunar surface were taken by the crew in the LM cabin. From Apollo 12, hammocks were supplied which could be put up across the living space making the rest period slightly more comfortable.

The environmental system was manufactured by Hamilton Standard Co. as was that of the command module, but did not supply hot water for the LM crew as that luxury was deemed unnecessary for its projected two day use. The system also provided the facility for connection of the crew's suits via umbilical cords to supply oxygen and cooling water during the landing and ascent phases of the flight. The ECS also provided a means of supplying the self contained EVA suits backpacks with its consumables of oxygen, water and battery power for use on the lunar surface. On later missions it allowed the suits consumable to be topped up between multiple EVAs.

Aft Equipment Bay

The aft, or rear equipment bay situated on the external wall at the rear of the crew compartment contained the majority of the LM's electronic equipment modules and communication equipment. Electronic module were fitted to heat sink rails and thermal control was achieved by glycol water mixture circulated through attachment rails and radiators in the same manner as that in the command module. It was controlled by the LM's ECS.

Direct communication between the LM and mission control during flight was through an S-Band transmitter-receiver using a single 26 inch diameter, steerable antenna, controlled by the guidance computer. This carried voice and television signals as well as spacecraft telemetry data. Two fixed, S-band aerials provided in-flight communication channels with the CM, which was backed up by two further VHF in-flight antennas. Once on the lunar surface communication between the crew and mission control was through two VHF channels relayed through the LM and the CM back to earth. On later flights an optional S-band antenna could be erected to provide direct communication with the crew on the lunar surface and mission control. A flashing strobe tracking light allowed visual contact between the two craft, visible over 75 miles.

The Ascent Propulsion System (APS) engine, manufactured by Bell Aerosystems, was a fixed 3,500 pound, constant thrust motor, which could accelerate the ascent stage from take off to a speed of 6,000 feet per second during its seven minute, once only firing. It used the same hypergolic fuels and pressurisation systems as the descent engine. Bell, also had their own development problems with the ascent engine. Erosion of the engine's ablative material in the combustion chamber's throat had been overcome early in the engine's development, but NASA found that Bell were using test criteria from their previous engine design which had been used in the Agena unmanned test vehicles and that the correct procedures had not been requested by Grumman.

To certify an engine that was to be used in a manned vehicle required more stringent testing and the inclusion of a 'bomb test'. To pass the test and 'man rate' the engine, the detonation of an explosive charge within the combustion chamber while running at full thrust was not expected to significantly interrupt its continued operation. When the bomb tests were carried out combustion instability was introduced from which the engine was not able to recover. Months of testing finally resulted in a solution with the introduction of a new fuel injector head manufactured by Rocketdyne.

Assembling and Launching 'The Stack'

The Saturn S-IC booster and the second stage S-II were both delivered to the Kennedy Space Centre (KSC) in Florida by sea-going barges from their respective manufacturers. The third stage S-IVB and the Apollo spacecraft were delivered by air in modified cargo aircraft irreverently referred to as the 'Pregnant Guppy'. The Apollo spacecraft modules were passed to the Manned Spacecraft Operations Building (MSOB) where they were connected up to Acceptance and Checkout Equipment (ACE) for final testing. The three individual stages of the Saturn launch vehicle were each passed directly to the Vehicle Assembly Building (VAB) for checkout and stacking.

Acceptance and Final Testing (MSOB)

The acceptance testing by NASA of the Apollo spacecraft from its manufacturers, was carried out under clinically clean conditions in the MSOB. Each spacecraft was linked to ACE computers and underwent tests that duplicated each phase of the mission it was to perform. They were also subjected to leak tests in vacuum chambers simulating altitudes of 200,000 feet. The complete mission was simulated several times in the course of the checkout and on successful completion the LM's ascent and descent stages were mated together for the final time and assembled in its adapter housing. The CSM was then assembled to top of the adapter housing, ready to be passed to the VAB for final mating with the Saturn launch vehicle.

The stacking process began with a mobile crawler that picked up a mobile launch pad and a Launch Umbilical Tower (LUT) and placed it in position in one of the four VAB assembly bays. The crawler was powered by two 2,750 horsepower diesel engines driving four caterpillar tracks, each link in the tracks weighing one ton. Each of the three stages were then mated vertically on the mobile launchpad and connected to the lower umbilical supply arms of the tower. The bottom S-1C stage was held in place by up to twelve pull-through bolts and four explosively operated, launch release clamps. All three stages and the instrument unit were connected to the umbilical tower for testing by the ACE computers prior to receiving the Apollo spacecraft. The Apollo spacecraft in its adapter housing was then mated to the top of the Saturn stack and connected to the LUT through which its systems could be again tested.

Roll Out to Launch Pad

On completion of construction, and when ready for launching, the complete Saturn/Apollo stack would begin its journey to the moon at a stately one mile per hour on its mobile crawler to the launch site three miles away. The VAB and the launch sites are connected by a crawlerway constructed to support the combined weight of the spacecraft, mobile launchpad, tower and crawler. During its journey and when climbing onto the raised launch apron, the upright stack was controlled by a level sensing device which kept it vertical to within ten seconds of arc.

At the end of its run the stack was centred over a 50ft deep flame trench by the crawler, where the Saturn's exhaust plume would be directed at a vee-shaped flame deflector in the trench to split and direct it sideways. As the S-IC engines ran up to full power, a water deluge of up to 50,000 gallons per minute was directed into the trench to cool the exhaust gasses and damp down their sonic vibration. Finally, an access gantry, the Mobile Service Structure (MSS), the height of the moonship stack, was brought up to the spacecraft, opposite the LUT, to provide all round access. The stack would then be connected up to the Launch Control Centre (LCC) computers.

Final tests and loading of Apollo spacecraft gasses, ordnance and cryogenic fuels preceded the commencement of the launch final countdown at T minus 28 hours. Removal of the MSS was completed at T minus 10 hours. After a built in hold of the countdown at T minus 9 hours, chilling and loading kerosene, liquid oxygen and hydrogen to the Saturn launch vehicle began, supplied from storage tanks adjacent to the site which were linked to the spacecraft through the LUT. The crew took their positions at T minus 2 hours 40 minutes and the firing command to the automatic sequencer for engine ignition given at T minus 3 minutes 6 seconds. Engine ignition commenced at T minus 8.9 seconds and lift-off at T minus 00:00. Some eight seconds after lift off, just after the stack cleared the launch tower, the responsibility for the conduct of the flight transferred from Kennedy Space Centre launch control to the Mission Control Centre (MCC) at Houston, Texas.