[6] The first four guidelines for the design of the advanced manned space vehicle deal with mission requirements, propulsion systems and design flight time capability. They are presented in figure 1.

1. The vehicle should be capable ultimately of manned lunar reconnaissance.

2. The lunar vehicle should be capable of earth orbital missions for initial evaluation and training. The reentry component of this vehicle should be capable of earth orbital missions in conjunction with space laboratories or space stations.

3. The multiman advanced space vehicle should be designed to be compatible with the Saturn and for the lunar mission it should weigh not more than 15,000 pounds including auxiliary propulsion and attaching structure.

4. The vehicle should be designed for a flight time capability, without resupply, of 14 days.

The first guideline states that the vehicle should be capable, ultimately, of manned, lunar reconnaissance. The justification for this step, in a technical sense, is that it is a logical intermediate step toward future goals of landing men on the moon and other planets. This mission will require solution of many of the problems associated with the manned, lunar landing mission. This is particularly true of the earth reentry and recovery phases of the flight. In addition, it is a mission which will require a considerable amount of trajectory control, consequently imposing rather severe requirements on the navigation and control system, and thereby effectively demonstrating our ability to navigate in space. A further significant consideration in selection of the manned lunar reconnaissance is that it is the ultimate manned mission compatible with our firmly planned booster program.

Numerous mission analyses involving trajectory studies will be required to determine the most desirable flight path for the vehicle. Figure 2 presents a typical lunar trajectory and return. It may be used to illustrate the initial approach to the mission that is being taken. This involves a flight path which would pass the moon at a distance on the order of several thousand miles. Such a flight path [7] can give a condition which will allow the vehicle to return to earth without the necessity of firing large rockets. These trajectories which will give this so called "free" or "safe" return will have to be identified by study, and their compatibility with launch insertion conditions and reentry conditions established.

To really accomplish lunar reconnaissance, prior to a manned landing, it is considered desirable to come much closer to the moon than several thousand miles. The order of 50 miles would appear to be a reasonable initial target for study purposes. To achieve this, we would propose to go into orbit about the moon. This procedure (as compared to a close pass by) has the advantage of allowing the decision to be made in flight after establishing the integrity of the system, and also allows loitering in the neighborhood of the moon for a time period sufficient for reconnaissance requirements.

Many other detailed trajectory studies will naturally be required. Several examples might be various midcourse abort trajectories, and the effects of the nonuniformity of the moon's shape on the vehicle trajectory while in the neighborhood of the moon.

While the ultimate capability of the proposed vehicle is manned lunar reconnaissance, an intermediate or parallel mission requirement must be considered in the design of the vehicle. The second guideline, therefore, states that "The lunar vehicle should be capable of earthorbital mission and that the reentry component of this vehicle should be capable of earth orbital missions in conjunction with space laboratories or space stations."

The justification for this requirement stems from two sources:

1. Before lunar missions are attempted, earth orbital flights will be required for lunar vehicle and crew evaluation' crew training and the development of operational techniques.

2. Concurrently' there is a need for a vehicle to be used in earth orbits with space stations for numerous research and development programs related to space flight.

Consequently, the reentry vehicle should be designed for use with both the lunar and orbiting space laboratories to avoid unnecessary duplication of vehicle development.

Figure 3 conceptually illustrates some of the possible combinations of components for use in the lunar and earth orbital flights. It is suggested that the lunar vehicle will consist of a reentry component [8] and a minimum "mission component" to stay within payload limitations, as shown on the figure. The "mission component" or "caboose" allows jettisoning of this portion of the vehicle prior to reentry' thereby not requiring protection from reentry heating. In addition' this arrangement appears to have good usable space characteristics and two compartments appear possibly desirable for emergency conditions. However, complete studies have not been carried to the point where it can definitely be said that the two component vehicle is superior from all standpoints as compared to using a single, larger vehicle, such as illustrated on the figure. Consequently' a detailed study needs to be made to determine the trade offs between using a one or two component lunar vehicle. Such a study should consider the optimum distribution of systems and supplies between the two components. On the right of figure 3 the reentry component is illustrated, which may be either of the two sizes shown on the left, with an earth orbital laboratory or space station. The large size of the space laboratory is allowable because of the increased earth orbital payload capability as compared to lunar payload capability. It is suggested that the first of these orbiting space laboratories might be made compatible with the early version of the Saturn and, therefore, should weigh 25,000 pounds or less including the reentry component.

The third guideline relates to the propulsion system and states that "The multiman advanced space vehicle should be designed to be compatible with the Saturn and for the lunar mission it should weigh not more than 15,000 pounds including auxiliary propulsion and attaching structure." The justification for this guideline is fairly straightforward in that the Saturn is the only firmly scheduled booster capable of the manned lunar mission.

Although not required for our guideline presentation, we have prepared, for background information' several figures which give some approximate Saturn characteristics.

The Saturn booster is being developed in several versions. The two versions with which we are most concerned are referred to as C-1 and C-2. Figure 4 presents sketches and characteristics of the booster. The booster overall lengths , without payloads, are 150 feet and 180 feet for the C-1 and C-2, respectively. It should be remembered that these boosters are in a development or preliminary design phase and all numbers given here are subject to modification as studies and development continue.

Saturn C-1. The first stage of C-1 is being designed and developed at the Marshall Space Flight Center in Huntsville, Alabama. The first booster is now on the test stand and testing is underway. This stage is [9] made up of a central tank and eight smaller tanks located around the central tank. Liquid oxygen is carried in the central tank and in four of the outer tanks. RP-1 is carried in the remaining four tanks. Control of the first stage is obtained by gimballing the four outer engines. The four central engines are fixed.

The second stage of C-1 will be developed by Douglas Aircraft Company. The stage uses liquid oxygen and hydrogen, as do all the upper stages of the Saturn vehicles. All four engines, which will be uprated Centaur engines, are gimballed for control.

The third stage is essentially the Centaur stage being developed by Convair/Astronautics.

Two stage C-1 vehicles may be available for research and development testing as early as December 1962. Vehicles qualified for manned use are estimated to be available in 1966. Ten flights are planned in this qualification program.

The C-1 version does not have sufficient payload capability for the manned circumlunar mission; however, it does have the capability of placing up to 25,OOO pounds of payload in earth orbits leading up to lunar missions.

Saturn C-2. The first stage of C-2 is similar to the first stage of C-1, except that its tankage may be reduced consistent with optimum propellant loading.

The second stage of C-2 is not used in C-1 but is a new stage. It will use four 200,000 pound thrust hydrogen oxygen engines. The contract for this engine is now being let. The stage contract has not been let; therefore, this is the least advanced of any of the Saturn stages.

The third stage of C-2 is the same as that used as the second stage of C-1.

C-2 vehicles will be available for research and development flight tests in 1965 and for manned tests in 1967. Four flights of this vehicle are planned for C-2 qualification, in addition to the ten planned for C-1.

Figures 5 and 6 present some launch trajectory parameters for the C-1 and C-2 vehicles. These trajectories have not been optimized for the missions considered and are presented here to indicate relative performance of stages, and the loadings to be expected in the way of dynamic pressure and axial accelerations.

[10] Figure 5 presents the variation of dynamic pressure, acceleration and velocity as a function of time for a launch trajectory into earth orbit using the C-1 vehicle. The maximum dynamic pressure is seen to be 775 pounds per square foot and occurs about midway through first-stage burning. For reference, the maximum dynamic pressure on the Mercury-Atlas flights is about 1,000 pounds per square foot. At first-stage burnout, at 210,000 feet altitude, the dynamic pressure has dropped to near zero.

The maximum axial acceleration of 6.5g occurs at first-stage burnout, and is several g less than experienced in the Mercury-Atlas flights.

The velocity contributions of the three stages are 9,000, 8,000, and 8,500 feet per second, respectively, to give the orbital velocity of 25,500 feet per second.

Figure 6 presents the variation of dynamic pressure, acceleration, and velocity as a function of time for an escape launch trajectory for the C-2 vehicle. Maximum dynamic pressure is seen to be 600 pounds per square foot as compared to 775 pounds for C-1. The dynamic pressure at separation is on the order of 400 pounds per square foot. The implications of separation dynamic pressures of this magnitude on the design of automatic abort-sensing systems, and on vehicle loading, in case of abort, will require careful consideration.

Maximum longitudinal acceleration is about 5g and occurs at secondstage burnout.

Velocity contributions of the three stages to escape velocity are 3,000, 15,000, and 18,000 feet per second, respectively, for the three stages.

The fourth guideline states that "The vehicle should be designed for a flight time capability, without resupply, of 14 days." In justification, it is to be noted that minimum time for a lunar circumnavigation mission is about 6 days. Taking into account that the total flight time may be affected by lunar perturbations, loiter times, or use of nonminimum time trajectories, considering the need for safety factors, and the desire for mission flexibility, it is believed that the vehicle should be designed for a minimum flight time of 14 days.

A particularly important study is required in regard to this capability in connection with the selection of a suitable type of power supply. It is roughly estimated that 400,000 watt-hours are required for the mission. Considerable study of storage batteries, fuel cells, auxiliary power units, and solar batteries has been made by a number of [11] people. The merits of these various systems or combinations must now be considered in the light of the particular requirements of this mission, with particular emphasis being placed on the reliability required for manned flight. Some items that should be considered in such a study include:

1. What percentage of the power units is desirable to place in the "caboose?"

2. Would the use of storage batteries both for power and as radiation shielding make this type of unit preferable?

3. Should redundancy for reliability be obtained by using two different types of systems or two of the same system?