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The Saturn/Apollo Stack - Lunar Module

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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

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 ''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 moon's 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.

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