Apollo Direct 2-Man

A direct lunar lander design of 1961, capable of being launched to the moon in a single Saturn V launch through use of a 75% scale 2-man Apollo command module.

Following the preliminary decision to proceed with the lunar orbit rendezvous technique for the Apollo lunar landing mission, there was one final effort to return to the simpler direct landing approach. Lunar orbit rendezvous would require a three-man Apollo capsule, in order for a crew of two to reach the lunar surface while the third crewmember tended the waiting Apollo in lunar orbit. But if the objective was to land two men on the moon, why use the three-man Apollo capsule? Why not use either the two-man Gemini capsule, or a reduced size two-man Apollo-shaped capsule to land directly on the moon? Such a spacecraft could be propelled toward the moon on a single launch of the Saturn C-5 rocket, just like the LOR version.

McDonnell, builders of the Mercury and Gemini spacecraft, were given the contract to make a study of the alternate approach. The result showed it was indeed feasible, at less cost, risk, complexity, and time, then the LOR 3-man Apollo. The only problem was that the existing work done on the North American three-crew Apollo would have to be scrapped, and either a modified Gemini or a new two-crew Apollo developed in its place. Wiesner, Kennedy's science adviser, was an enthusiastic about the approach. NASA, North American, and the competitors for the lunar module contract were distinctly less interested. Webb, the NASA Administrator, finally got the idea spiked once and for all.

Major ground rules for the study were:

• Launch vehicle was the Saturn C-5, injecting 40,800 kg to a lunar transfer orbit.
• Mission duration was eight days (two and one-half days flight time to moon, one day on lunar surface plus one day contingency, two and one-half days return time plus one day contingency) plus seven days post landing (one day habitable environment plus six days survivable environment ).
• Both cryogenic and noncryogenic propulsion systems were to be considered. The final configuration used a LOX/LH2 RL10-powered stage for the major lunar crasher stage, and N2O4/MMH storable propellants for the lunar landing and ascent stages.
• Atmosphere was 5 psi normal, 3.5 psi emergency, 100% oxygen.
• First operational flight during the first half of calendar year 1967 was assumed.

The command module was a 33 decree blunt cone with the same geometry as the Apollo 3-man capsule then in development. But it was 25% smaller, with a 3.18 m diameter base. The module has a hypersonic L/D of 0.45, achieved by offset center of gravity. The two crewmen entered and exited the module through a quick opening hatch in the side wall. A tower mounted launch escape system was attached to the conical end of the command module.

Weight summary was as follows:

• Command Module: 2337 kg
• Service Module Equipment: 520 kg
• Spacecraft Weight Margin: 652 kg
• Service Module: 10,104 kg
• Terminal Landing Module: 2735 kg
• Retrograde Module: 24,393 kg
• Launch Escape System: 1,179 kg
• Landing Gear Fairing: 634 kg
• Gross Weight at Launch: 42,556 kg
• Less (1% effective) Jettisonable Items: 1740 kg
• Effective Launch Weight: 40,816 kg

The guidance and navigation system for Two-Man Apollo was essentially the same as the system developed by MIT for the Three-Man Apollo. It consisted of an inertial system for attitude memory and acceleration measurements, navigation sensors (optical, radar, and DSIF), a computer for computation of guidance and display signals for nominal and abort trajectories, and electronic interface equipment required to drive appropriate displays. Total weight was 176 kg, compared with 200 kg for the Three- Man Apollo.

The environmental control system was identical to the equipment in the Gemini spacecraft with the exception of the volume differences dictated by the requirements for expendables. Normal mission power was supplied from six fuel cell stacks located in the service module. Silver-zinc batteries in the command module provided re-entry and recovery power plus back-up for the fuel cells, as required. Additional silver-zinc batteries provided redundant isolated sources for squibs, relays and other transient producing equipment. Both the fuel cells and the batteries were identical to Gemini equipment. A detailed electrical load analysis indicated that the mission requirements were 900 watts average. Sufficient fuel was provided for the full 8 day mission, two days of which were contingency.

Conventional spacecraft structures were employed in all modules, following the proven materials and concepts demonstrated in the Mercury and Gemini designs. Primary structure of each module consisted of a semimonocoque shell with reinforcements around cut-outs and fittings to distribute localized loads. Titanium was used as the basic shell material in all modules except the service module where beryllium sheet was used for the structural radiator shell.

Re-entry heat protection was conservatively designed for a controlled 17,000 km range, polar re-entry. The ablative material was M.A.C. Thermorad Shield S-3 elastomeric composite, which was being used for the Gemini blunt face heat shield. A 1.15 factor was applied to the predicted heating rates. Nominal material physical properties were used, determined from Project Gemini testing. Total ablative material design weight was 259 kg.

• Apollo Direct CM.  Manufacturer's Designation: McDonnell. Class: Manned. Type: Spacecraft Module. Nation: USA. Agency: NASA. Manufacturer: McDonnell.

  Conventional spacecraft structures were employed, following the proven materials and concepts demonstrated in the Mercury and Gemini designs. Primary structure consisted of a semimonocoque shell with reinforcements around cut-outs and fittings to distribute localized loads. Titanium was used as the basic shell material. Re-entry heat protection was conservatively designed for a controlled 17,000 km range, polar re-entry. The ablative material was M.A.C. Thermorad Shield S-3 elastomeric composite, which was being used for the Gemini blunt face heat shield. A 1.15 factor was applied to the predicted heating rates. Nominal material physical properties were used, determined from Project Gemini testing. Total ablative material design weight was 259 kg.

  Crew Size: 2. Length: 2.67 m (8.75 ft). Basic Diameter: 3.18 m (10.43 ft). Mass: 2,337 kg (5,152 lb). L/D Hypersonic: 0.45. Electrical System: Batteries.

• Apollo Direct SM.  Manufacturer's Designation: McDonnell. Class: Manned. Type: Spacecraft Module. Nation: USA. Agency: NASA. Manufacturer: McDonnell.

  The Service Module housed the fuel cells, environmental control, and other major equipment items required for the mission. It also provided the propulsion requirements for lunar launch through transearth midcourse corrections. The same basic engine would used as on the terminal landing module with the throttling feature removed and a radiation cooled skirt added to provide an expansion ratio of 40:1.

  Length: 2.60 m (8.50 ft). Basic Diameter: 2.85 m (9.35 ft). Maximum Diameter: 5.89 m (19.32 ft). Mass: 10,624 kg (23,421 lb). Payload: 520 kg (1,140 lb). Main Engine Propellants: N2O4/MMH. Main Engine Propellants: 8,840 kg (19,480 lb). Electrical System: Fuel Cells.

• Apollo Direct TLM.  Manufacturer's Designation: McDonnell. Class: Manned. Type: Spacecraft Module. Nation: USA. Agency: NASA. Manufacturer: McDonnell.

  Final letdown, translation hover and landing on the lunar surface from 1800 m above the surface was performed by the terminal landing module. Engine thrust could be throttled down to 1546 kgf. The engine was presumed to be based on a reasonable development of any one of three on-going engine programs for Apollo of the time: Aerojet General's 3600 kgf thrust Transtage engine; Reaction Motors' Model TD-294 4500 kgf thrust LEM research engine; or the 4500 kgf thrust engine under investigation by Rocketdyne for possible use in the LEM. The engine in the terminal landing module was ablatively cooled, throttleable, and had an expansion ratio of 20:1.

  Length: 1.98 m (6.49 ft). Basic Diameter: 5.89 m (19.32 ft). Maximum Diameter: 6.61 m (21.68 ft). Mass: 2,773 kg (6,113 lb). Recovery Equipment: 336 kg (740 lb). Main Engine Thrust: 50.121 kN (11,268 lbf). Main Engine Propellants: N2O4/MMH. Main Engine Propellants: 1,533 kg (3,379 lb).

• Apollo Direct RM.  Manufacturer's Designation: McDonnell. Class: Manned. Type: Spacecraft Module. Nation: USA. Agency: NASA. Manufacturer: McDonnell.

  The retrograde module supplied the velocity increments required during the translunar portion of the mission up to a staging point approximately 1800 m above the lunar surface. Engine thrust could be throttled down to 1201 kgf.

  Length: 5.00 m (16.40 ft). Basic Diameter: 6.61 m (21.68 ft). Maximum Diameter: 6.61 m (21.68 ft). Mass: 24,393 kg (53,777 lb). Main Engine: RL10A-3. Main Engine Thrust: 117.856 kN (26,495 lbf). Main Engine Propellants: LH2/Lox. Main Engine Propellants: 21,560 kg (47,530 lb).

• NASA Report, Direct flight Apollo Study Briefing- October 1962, . Accessed at: http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19790077943_1979077943.pdf. -right click to download pdfs!

• NASA Report, Direct Flight Apollo Study. Volume 1: Two Man Apollo Spacecraft - October 1962, . Accessed at: http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19730064943_1973064943.pdf. -right click to download pdfs!