Moon Lander: How We Developed the Apollo Lunar Module (15 page)

The LM program in Project Christmas Present included ten flight LMs, the first two of which were unmanned, and six LM test articles for ground tests, as follows: LTA-2, launch vibration tests at Huntsville; LTA-10, SLA fit and mass model tests at North American-Tulsa; LTA-1, “house spacecraft” at Bethpage for electronics tests and support of fabrication, assembly, and checkout; LTA-8, thermal vacuum tests at Houston; and LTA-3 and LTA-5, structural and vibration tests at Bethpage. The ground-test program also contained boiler-plate and flight-weight propulsion test rigs for fluids tests in the cold flow facility at Bethpage and hot rocket firings at White Sands. The first unmanned LM flight was scheduled for late 1967, and the ground test articles were planned for use during 1966 and 1967.

Grumman Leads Mission Planning

As our LM designs took shape we encountered questions on LM requirements that could only be answered by a better understanding of how the lunar landing mission would be conducted. The design of the crew compartment, for example, was greatly influenced by the crew’s activities on the lunar surface. When would the spacesuits be worn? How many times would they be doffed and donned in the LM cabin, and how much volume would this require? What about the return of lunar samples; size, weight, contents? One set of questions generated others. Determining the required capacity and duty cycle of the EPS called for detailed knowledge of the mission time line; accounting hour-by-hour for what mission activities were in progress and what equipment was turned on. The same was true of thermal analysis, determining the heat loads used to size the ECS. Communications system requirements, duty cycles, and antenna positioning and usage could not be finalized without a detailed mission plan, nor could any of the LM systems. I realized that we must have a mission baseline for our design.

Rathke and I proclaimed the need for a design reference mission (DRM) to give all Apollo contractors a basis for finalizing their system design requirements. We asked Tom Barnes to take our existing lunar-orbit rendezvous studies and expand their mission definition. We also began talking up the idea informally at Apollo program meetings with NASA, NAA, and MIT.

In September 1963, shortly after the approval of a definitive contract with North American Aviation, Joseph F. Shea replaced Charles Frick as Apollo spacecraft program manager in Houston.
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Joe Shea was brought into NASA’s newly formed Office of Manned Space Flight in November 1961 from Space Technology Laboratories by NASA’s administrator, James Webb, and D. Brainerd Holmes, associate administrator for Manned Space Flight.
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Known from the Titan and Minuteman ballistic missile programs as a brilliant aerospace systems engineer, he was made Holmes’s deputy and given the challenge of evaluating the competing lunar mission modes. Shea was an articulate man of overpowering intellect, a skilled debater, persuasive in argument and a powerful program leader. He was tall and handsome, with an athletic build and dark Irish good looks—jutting jaw, fair complexion, prominent black eyebrows, and black crew-cut hair. He was dangerously dynamic, so likely to prevail in most arguments that the entire Apollo program depended upon his wisdom and good judgment. Although he could witheringly destroy an opponent’s arguments as capably as any trial lawyer, he also had a wonderful sense of humor and enlivened many a meeting with his witty, inventive puns.

When I mentioned to Shea our need for better definition of the lunar mission to pin down LM design requirements, he threw the ball back to me by recommending that Grumman lead a mission study with participation by the other Apollo contractors. Thus was born the Apollo Mission Planning Task Force (AMPTF) in January 1964. With Barnes in charge, the AMPTF set up shop in one of the large Apollo conference rooms in Plant 25 and was joined by team members from NAA, MIT, and NASA-Houston. Tom Barnes was a great team leader. Friendly, constructive, and totally without ego or institutional bias, he inspired confidence and cooperation from the entire task force. Barnes was a talented systems engineer who explored problems relentlessly, asking key “what if” questions that sometimes led to new ways of defining or resolving things. He did it in such an easygoing but provocative manner that others were stimulated to new insights and contributions.

The task force started by defining the basic mission objectives: “Land two astronauts and scientific equipment on the near-Earth-side surface of the Moon and return them safely to Earth.” A second objective was to carry at least 250 pounds of scientific equipment to be set up on the Moon and to bring back 100 pounds of lunar soil and rocks.
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The AMPTF created a detailed description and analysis of all flight mission activities from liftoff to splashdown and recovery—the DRM. They also investigated possible failure modes and contingencies to determine their effect on mission planning and on spacecraft design requirements.

Four months of intensive mission planning and analysis took place with dozens of engineers from NASA and the contractors participating. To make possible precise launch and trajectory calculations using the exact relative positions of Earth, Moon, and spacecraft and figuring the rocket firings necessary
to execute mission maneuvers and flight path corrections, it was necessary to choose a specific date for the DRM. The team selected 6 May 1968 for liftoff. Using the same minute-by-minute crew time line planning technique that had been developed on Projects Mercury and Gemini, the AMPTF extended the detailed flight plan to cover three astronauts and two spacecraft (CSM and LM) that for part of the mission functioned independently. Actions required of the crew, the spacecraft, and the ground network to perform the mission were documented in the DRM time line, and trajectory calculations and error analyses were performed to establish system performance and accuracy requirements. The result was the most complete prelaunch mission planning yet attempted for Apollo, providing a good basis for further development of design requirements, operational ground rules, and mission plans.

The DRM clarified the docking requirements for both the CSM and LM. Initial docking and extraction of the LM from the SLA would be carried out by the crew from their couches in the CM. The resulting connection would provide a rigid pressurized tunnel permitting crew access between both spacecraft. Upon return from the Moon, the rendezvous maneuver would be performed by the two-man LM crew, with CM-active backup available, but the docking would be done by the lone crewman in the CM in the same manner as the initial docking. Only the upper LM hatch was required for docking, whereas the forward LM hatch was needed for lunar surface egress.

The DRM became a treasure trove of information for contractor engineers seeking firm requirements to which to design their systems and components. At Grumman we set up a formal process to tabulate the requirements specified or implied by the DRM and compare them with the design specifications for the LM spacecraft and its systems and components; correcting the design specs where necessary to bring them into conformance. This assured that the spacecraft we were designing and subcontracting would be able to perform its overall intended function of lunar landing and return.

The planners looked for failure modes at each step of the mission and sketched out recovery plans where possible. They also determined the accuracy required in critical mission phases. For example, the midcourse trajectory corrections on the way to the Moon had to be accurate within three or four feet per second or else the spacecraft would crash into the surface. Upon reentry the CM must hit the outer edge of the Earth’s atmosphere within a flight path angle “window” of two degrees. Too steep an angle would result in a rapid plunge into the atmosphere, burning up the spacecraft like a meteor, while too shallow an approach angle would skip the CM off the top of the atmosphere and send it on an eternal orbit of the Sun.

One major result of the AMPTF contingency planning was the identification of the “LM Lifeboat” mission. While postulating the effect of various CSM failures on the outbound leg of the mission, the planners realized that a number of them could be countered by using the LM as a lifeboat and utilizing
its propulsion, guidance and control, life-support, and other systems to return the crew to the vicinity of the Earth’s atmosphere for reentry in the CSM. To provide this rescue capability, some of the LM consumables, such as oxygen, water, and electrical power, would have to be increased by 10 to 15 percent above that needed to perform the basic mission. Because LM then existed only on paper, we decided to make the tanks that much larger. At a later date it could be decided whether to actually load the additional consumables into them. Six years after it first appeared in the AMPTF’s report, this vital crew rescue mode was dramatically utilized on Apollo 13.

Loose Cannon

As we performed comparative analyses on the LM systems we came to doubt the reliability estimates for the MIT guidance, navigation, and control system that was provided to both the LM and the CM. We arrived at this conclusion while preparing our own estimates of the reliability of the backup abort guidance system, for which Grumman was responsible. This led us to challenge MIT’s GNC reliability estimates, with disastrous results for Grumman.

Our GNC experts came to believe, and they convinced me, that MIT’s GNC had a factor of one hundred lower reliability than MIT claimed, an opinion based mainly upon Grumman’s interpretation of summary mean-time-between-failure (MTBF) data for guidance system components on the Polaris, Titan, and Minuteman ballistic missile programs, as published in a GE report. It was instigated by a former Honeywell reliability engineer working for Grumman who may have had a hidden agenda. (MIT had beaten Honeywell in the competition for the Apollo GNC and said engineer was on the Honeywell proposal team.) After Grumman published its conclusions in a report to the Apollo program office, MIT and NASA hit the roof.

Joe Shea convened a meeting of all interested parties in early January 1964 to find the truth and punish the guilty. We gathered in the well-appointed Apollo program conference room in Houston. About thirty NASA, MIT, air force, and Bellcom GNC experts attended, led by Jim Elms, representing Bob Gilruth, MSC director, and Joe Shea, Apollo spacecraft program manager. Gavin, Rathke, Whitaker, and I sat opposite this stone-faced group. Shea warned everyone that the meeting was being tape-recorded.

Elms welcomed everyone and said the meeting would be a technical discussion to resolve questions concerning Apollo guidance system reliability. He turned the meeting over to Shea. Shea glowered under his black eyebrows and hunched forward toward his microphone, his hands knotted tightly together on the desk, his gold MIT ring visible:

Gentlemen, we have a serious problem. The problem is that Grumman believes the MIT guidance system is two orders of magnitude inferior to other available
systems, and that the Apollo program is being jeopardized by this choice. The issue centers upon the evaluation of the data used in drawing this conclusion and establishing its validity.

I intend to force a black-and-white conclusion as a result of this meeting. Either there is or there is not a significant basic difference in the inherent reliability of the MIT system and other comparative data. Someone, Grumman or MIT, will have to leave this meeting admitting he was wrong—mea culpa, mea maxima culpa.

Shea beat his breast for emphasis. He explained the points of interest in reliability data assessment, lecturing us like schoolboys. The sources, accuracy and time frame of the data, the relevance of failures included, and the type of program all must be interpreted to select data that is comparable to the Apollo system.

Whitaker took the stand to present Grumman’s case.

He said that our evaluation of the inertial guidance system reliability data that we had been able to obtain predicted 894 failures per million hours, versus 10 failures per million hours estimated by MIT. This puzzled us and caused us to consider other design approaches for the LM abort guidance system, such as strapped-down inertial components instead of a gimbaled platform. In discussions with MIT and with guidance system manufacturers we had not been able to resolve this discrepancy. If Grumman had misinterpreted the data or overlooked other relevant data sources, we were open to correction.

Whitaker reviewed the Polaris data that Grumman used. Most of it had come from Minneapolis Honeywell, which manufactured both a MIG gyroscope (gyro) that they designed and the MIT-designed IRIG gyro. Honeywell’s numbers showed a much lower failure rate for the MIG gyro.

Next Whitaker turned to Titan program data. Grumman had only found a small amount of Titan data, and it showed high failure rates. Because the Titan GNC design was similar to MIT’s Apollo system, we used this data in calculating our failure rate.

“Didn’t you think there must be a lot more Titan data available than that?” bristled Shea, who led the development of the Titan II guidance system at General Motor’s AC Electronics Division. “You’re implying that all the Titan data you couldn’t find also showed high failure rates. Well you’re wrong, as you’ll soon find out.”

Things got worse. When Whitaker showed our use of Minuteman data, the air force said that none of it was applicable because it all related to the percentage of the missile force available on thirty seconds warning—it did not relate to predicting design reliability. Whitaker sat down dejectedly.

Dave Hoag gave MIT’s rebuttal. Starting with Grumman’s number, 894 failures per million hours, he patiently showed how that number dropped as each incorrectly included or omitted set of data was corrected. Grumman’s data was based upon 130 reported failures during 158,000 hours of operation,
during the time period covered by the GE report. MIT had more complete results from GE, listing total failures and correcting errors in the report, and these showed 118 failures during 380,000 hours. Eliminating failures of rebuilt units and of designs subsequently scrapped dropped the relevant failures to 60, yielding a failure rate of 196 per million hours. From there on, differences between the design details of the Polaris, Titan, and Minuteman systems from Apollo had to be accounted for. Step-by-step Hoag explained the rationale for these corrections and showed the resultant drop in the failure rates. After two hours of logical explanation he had walked the MIT failure rate down to 20 failures per million hours. He admitted that their published estimate of ten per million was insupportable—attributed to program manager’s optimism.