The original Apollo spacecraft Statement of Work of July 28 had been substantially expanded, including a single-engine service module propulsion system using Earth-storable, hypergolic propellants.

The requirements for the spacecraft navigation and guidance system were defined:

Control of translunar injection of the spacecraft and monitoring capability of injection guidance to the crew both for direct ascent and for injection from an earth parking orbit.

Data and computation for mission abort capability en route to the moon and for guidance to a point from which a safe lunar landing could be attempted.

Guidance of the command module to a preselected earth landing site after safe reentry.

Guidance for establishing lunar orbit and making lunar landings; mission abort capability from the lunar landing maneuver.

Control of launch from the lunar surface into transearth trajectory by both direct ascent and from lunar parking orbit.

Rendezvous in earth orbit between the spacecraft and space laboratory module or other space vehicle.

Components of the navigation and guidance system now clearly identified were:

Inertial platform

Space sextant

Computer

Controls and displays

Electronics assembly

Chart and star catalog

Range or velocity measuring equipment for terminal control in rendezvous and lunar landing

Backup inertial components for emergency operation

The stabilization and control system requirements were revised:

Roll control as well as flight path control during the thrusting period of atmospheric abort and stability augmentation after launch escape system separation

Stabilization of the spacecraft and the lunar injection configuration while in earth parking orbit

Rendezvous and docking with the space laboratory module or other space vehicle

Attitude control and hovering for lunar landings and launchings and for entering and leaving lunar orbit

Basic components of the stabilization and control system were defined:

Attitude reference

Rate sensors

Control electronics assembly

Manual controls

Attitude and rate displays

Power supplies

The new system would be capable of:

Abort propulsion after jettison of the launch escape system

All major velocity increments and midcourse velocity corrections for missions prior to the lunar landing attempt

Lunar launch propulsion and transearth midcourse velocity correction.

The reaction control systems for the command and service modules would now each consist of two independent system, both capable of meeting the total torque and propellant requirements. The fuel would be monomethylhydrazine and the oxidizer would be a mixture of nitrogen tetroxide and nitrous oxide.

The parachute system for the earth landing configuration was revised to include two FIST-type drogue parachutes deployed by mortars.

The command module structure was specified: a ring-reinforced, single- thickness aluminum shell pressure vessel separated from the outer support structure of relatively rigid brazed or welded sandwich construction. The ablative heatshield would be bonded to this outer structure.

Service module structure was also detailed: an aluminum honeycomb sandwich shell compatible with noise and buffet and with meteoroid requirements. The structural continuity would have to be maintained with adjoining modules and be compatible with the overall bending stiffness requirements of the launch vehicle.

The duties of the three Apollo crewmen were delineated :

Commander

Control of the spacecraft in manual or automatic mode in all phases of the mission

Selection, implementation, and monitoring of the navigation and guidance modes

Monitoring and control of key areas of all systems during time-critical periods

Station in the left or center couch

Co-Pilot

Second in command of the spacecraft

Support of the pilot as alternative pilot or navigator

Monitoring of certain key parameters of the spacecraft and propulsion systems during critical mission phases

Station in the left or center couch

Systems Engineer

Responsibility for all systems and their operation

Primary monitor of propulsion systems during critical mission phases

Responsibility for systems placed on board primarily for evaluation for later Apollo spacecraft

Station in the right-hand couch.

One crew member would stand watch during noncritical mission phases at either of the two primary duty stations. Areas for taking navigation fixes, performing maintenance, food preparation, and certain scientific observations could be separate from primary duty stations. Arrangements of displays and controls would reflect the duties of each crewman. They would be so arranged that one crewman could return the spacecraft safely to earth. All crewmen would be cross-trained so that each could assume the others' duties.

Radiation shielding for the crew would be provided by the mass of the spacecraft modules.

A description of crew equipment was added:

The couch for each crewman would give full body and head support during all normal and emergency acceleration conditions. It would be adjustable to permit changes in body and leg angles and would be so constructed as to allow crewmen to interchange positions and to accommodate a crewman wearing a back or seat parachute. A restraint system would be provided with each couch for adequate restraint during all flight phases. Each support and restraint system would furnish vibration attenuation beyond that needed to maintain general spacecraft integrity. This system would keep crew vibration loads within tolerance limits and also enable the crew to exercise necessary control and monitoring functions.

The spacecraft would be equipped with toilet facilities which would include means for disinfecting the human waste sufficiently to render it harmless and unobjectionable to the crew. Personal hygiene needs, such as shaving, the handling of nonhuman waste, and the control of infectious germs would be provided for.

Food would be dehydrated, freeze-dried, or of a similar type that could be reconstituted with water if necessary. Heating and chilling of the foods would be required. The primary source of potable water would be the fuel cells. In addition, sufficient water would have to be on board at launch for use during the 72-hour landing requirement in case of early abort. Urine would not have to be recycled for potable water.

Emergency equipment would include:

Personal parachutes

Post-landing survival equipment:

one three-man liferaft, food, location aids, first aid supplies, and accessories to support the crew outside the spacecraft for three days in any emergency landing area. In addition, a three-day water supply would be removed from the spacecraft after landing; provision for purifying a three-day supply of sea water would be included.

Medical instrumentation would be used to monitor the crew during all flights, especially during stressful periods of early flights, and for special experiments to be performed in the space laboratory module and during extravehicular activity and lunar exploration. Each crewman would carry a radiation dosimeter.

The environmental control system would comprise two air loops, a gas supply system, and a thermal control system.

One air loop would supply the conditioned atmosphere to the cabin or pressure suits. The other would remove sensible heat and provide cabin ventilation during all phases of the mission including postlanding.

The primary gas supply would be stored in the service module as supercritical cryogenics. The supply would be 50 percent excess capacity over that required for normal metabolic needs, two complete cabin repressurization, a minimum of 18 airlock operations, and leakage. Recharging of self-contained extravehicular suit support systems would be possible.

Thermal control would be achieved by absorbing heat with a circulating coolant and rejecting this heat from a space radiator. During certain mission modes, other cooling systems would supplement or relieve the primary system.

Water collected from the separator and the fuel cells would be stored separately in positive expulsion tanks. Manual closures, filters, and relief valves would be used where needed as safety devices.

Metabolic requirements for the environmental control system were:

Total cabin pressure (oxygen and nitrogen mixture): 7 +/- 0.2 psia

Relative humidity: 40 to 70 percent

Partial pressure carbon dioxide - maximum 7.6 mm Hg

Temperature: 75 degrees F +/- 5 degrees F

The major components of the electrical power system were described more fully:

Three nonregenerative hydrogen-oxygen fuel cell modules characterized by low pressure, intermediate temperature, Bacon-type, utilizing porous nickel, unactivated electrodes, and aqueous potassium as the electrolyte

Mechanical accessories, including control components, reactant tankage, piping, etc.

Three silver-zinc primary batteries, each having a normal 28-volt output and a minimum capacity of 3,000 watt-hours (per battery) when discharged at the ten-hour rate at 80 degrees F

A display and control panel, sufficient to monitor the operation and status of the system and for distribution of generated power to electrical loads as required

The fuel cell modules and control, tanks (empty), radiators, heat exchangers, piping, valves, total reactants plus reserves would be located in the service module. The silver-zinc batteries anti electrical power distribution and controls would be placed in the command module.

Under normal operation, the entire electrical power requirements would be supplied by the three fuel cell modules operating in parallel. The primary storage batteries would be maintained fully charged under this condition of operation.

If one fuel cell module failed, the unit involved would automatically be electrically and mechanically isolated from the system and the entire electrical load assumed by the two remaining fuel cells. The primary batteries would remain fully charged.

If two fuel cell modules failed, they would be isolated from the system and the spacecraft electrical loads would immediately be reduced by the crew and manually programmed to hold within the generating capacities of the remaining fuel cell.

At reentry, the fuel cell modules and accessories would be jettisoned. All subsequent electrical power requirements would be provided by the primary storage batteries.

Each fuel cell module would have a normal capacity of 1,200 watts at an output voltage of 28 volts and a current density conservatively assigned so that 50 percent overloads could be continuously supplied. The normal fuel cell operating pressure and temperature would be about 60 psia and 425 degrees F to 500 degrees F respectively. Under normal conditions of operation, the specific fuel (hydrogen and oxygen) consumption should not exceed a total of 0.9 lb/kw-hr.

Self-sustaining operation within the fuel cell module should begin at a temperature of about 275 degrees F. A detection system would be provided with each fuel cell module to prevent contamination of the collected potable water supply.

The degree of redundancy provided for mechanical and electrical accessory equipment would be 100 percent.

The distribution portion of the electrical power system would contain all necessary buses, wiring protective devices, and switching and regulating equipment.

Sufficient tankage would be supplied to store all reactants required by the fuel cell modules and environmental controls for a 14-day mission. The reactants would be stored supercritically at cryogenic temperatures and the tankage would consist of two equal volume storage vessels for each reactant. The main oxygen and nitrogen storage would supply both the environmental control system and the fuel cells.

The communication and instrumentation system was further detailed:

The equipment was to be constructed to facilitate maintenance by ground personnel and by the crew and to be as nearly self-contained as possible to facilitate removal from the spacecraft. Flexibility for incorporation of future additions or modifications would be stressed throughout the design. A patch and programming panel would be included which would permit the routing of signal inputs from sensors to any selected signal conditioner and from this te any desired commutator channel. Panel design would provide the capability of "repatching" during a mission. The equipment and system should be capable of sustained undegraded operation with supply voltage variation of +15 percent to -20 percent of the normal bus voltage.

A circuit quality analysis for each radiating electrical system would be required to show exactly how ranging, telemetry, voice, and television data modulated all transmitters with which they were used.

The equipment and associated documentation would be engineered for comprehensive and logical fault tracing.

Components of the communication subsystem would include:

Voice communication

Telemetry

Tracking transponders

Television

Radio recovery aids

Antenna subsystems

Radar altimeter (if required by the guidance system)

The instrumentation system would be required to detect, measure, and display all parameters needed by the crew for monitoring and evaluating the integrity and environment of the spacecraft and performance of the spacecraft systems.

Data would be transmitted to ground stations for assessment of spacecraft performance and for failure analysis. Information needed for abort decisions and aid in the selection of lunar landing sites would also be provided. The mission would be documented through photography and recording.

Included in the components of the instrumentation system were:

Sensors

Data disposition

Tape recorders

Panel display indicators

Calibration

Clock

Telescope

Cameras

The propulsion system for the lunar landing module would now comprise a composite propulsion system: multiple lunar retrograde engines for the gross velocity increments required for lunar orbiting and lunar landing; and a lunar landing engine for velocity vector control, midcourse velocity control, and the lunar hover and touchdown maneuver. The lunar retrograde engines would use liquid-oxygen and liquid-hydrogen propellants. The single lunar landing engine would require the same type of propellant, would be throttleable over a ratio of +/- 50 percent about the normal value, and would be capable of multiple starts within the design operating life of the engine.

No additions or changes had been made in the space laboratory module systems description.

Overall control of all Apollo support elements throughout all phases of a mission would be exercised by the Mission Control Center. Up to the time of liftoff, mission launch activities would be conducted from the launch control center at Cape Canaveral. Remote stations would be used to support near-earth and lunar flights and track the command module during reentry.

Five major phases of a development and test plan were identified:

1. Design information and development tests
2. Qualification, reliability, and integration tests
3. Major ground tests
4. Major development flight tests
5. Flight missions.