[15] The following presentation is primarily concerned with the vehicle's return to the earth's surface at the termination of either a normal mission or an aborted mission. There are five guidelines in this area and they are shown in figure 1.
The first guideline requires that the vehicle should have the capability of safe crew recovery from aborted missions at any speed up to the maximum velocity, and that this capability be independent of the use of the launch propulsion system. It would follow that safe control of the vehicle subsequent to aborting the mission is thereby required. This control may be exerted by the auxiliary rocket propulsion (maneuvering rockets) or the aerodynamic L/D capabilities of the configuration.
The second guideline requires that the vehicle should be capable of a satisfactory landing on both land and water and should have the capability of avoiding local hazards in the landing area. This requirement was predicated upon the following considerations:
(a) Emergency conditions or navigation errors that force a landing on either land or water.
(b) Accessibility for recovery is an important consideration and the relative superiority of land versus water landing will depend on local conditions as well as other considerations. From our Mercury experiences, we have found that weather conditions which may exist over the earth's surface during a mission play an important part in determining the landing capability of the vehicle. Therefore, we feel that the vehicle's capability should include landing in a 30 knot wind, and that the vehicle should be water tight and have adequate sea keeping qualities under sea state 4 conditions (10-12-foot waves)
The third guideline requires that the vehicle should have the normal capability of landing at one of several previously designated ground surface locations, each approximately 10 square miles in area. In order to meet this requirement, studies are needed to assess the value of impulse maneuvers, guidance quality, and aerodynamic L/D during return from the lunar mission. It is desired, of course, that a suitable tradeoff between these items be arrived at to facilitate this performance. It might be mentioned that this requirement for point landing is far less severe for the orbital mission than for the lunar return situation. [16] In the fourth guideline, it is required that the vehicle be designed for crew survival for at least 72 hours after landing. Due to the random nature of possible emergency maneuvers, it will be impossible to provide sufficient recovery forces to cover all possible landing locations. The above requirement will permit mobilization of normally existing facilities and adequate time for a safe recovery. The vehicle, of course, should also be equipped with adequate location devices that will perform suitably in the various regions of the earth.
Finally, in the last guideline, it is required that adequate auxiliary propulsion should be provided for the guidance maneuvers required in the assigned missions and to effect a safe return in event of launch emergencies. The guidance accuracies and capabilities should be studied in order to determine the auxiliary propulsion requirements. Sufficient reserve propulsion should be included to accommodate corrections for maximum guidance errors. A single system may suffice for both the guidance maneuvers and the escape propulsion requirements.
The remainder of the presentation will provide additional discussion and material associated with the above guidelines.
First, let us consider the escape problem which will exist for this new mission, as shown in figure 2. This figure shows the various escape regimes that will exist for a launching to escape velocity. These regimes are presented as follows:
Launching period (including time on the pad prior to launch). In this regime, we are primarily concerned with providing sufficient altitude after escape from the propulsion system in order to satisfactorily deploy the landing parachute or, in the case of a winged vehicle, the escape maneuver must provide sufficient velocity to attain flying speed.
Atmospheric flight. This portion of the launch trajectory is within the atmosphere; here the problem of escape is associated with aerodynamic loads induced on the manned vehicle during and after the escape maneuver. It is important that the escape propulsion system has sufficient power to overcome the drag of the vehicle and pull it away from the booster system. It is also important that attention be paid to the stability characteristics of the vehicle during the escape maneuver.
Suborbital space flight. This regime starts at the time when the vehicle leaves the atmosphere and extends up to the time when orbital velocity is attained. In this case, the concern is with the subsequent reentry following an escape maneuver. It is important that the launch [17] trajectory and the escape maneuver be made in a manner that reentry deceleration loads will not exceed the tolerance of the crew.
We are rather familiar with the problems associated with these first three escape regimes as they all exist in the Mercury launch trajectory. It might be mentioned at this time that all of the critical heating and air loads associated with the design of the Mercury capsule were generated in analysis of the maneuvers resulting in escape, or subsequent to escape, during the reentry. Therefore, it is very important that close attention be paid to these escape regimes inasmuch as they most likely also provide the basis for all loads criteria for the advanced vehicle.
Superorbital space flight. The final escape regime extends from the time at which orbital velocity is reached up until the actual burnout of the propulsion system where escape velocity is achieved. In this regime, it is necessary to provide an adequate auxiliary propulsion system which can return the capsule to the surface of the earth in a reasonably short time. In fact, it is desired that descent to the earth's surface commence immediately with the escape maneuver. This requirement results from the desire to avoid long dwell periods in either the lower or upper radiation belt which may otherwise result if a propulsion failure left the vehicle in a highly elliptical orbit.
Superorbital escape trajectories. The most reasonable way to dissipate the energy that exists in the space vehicle is to use atmospheric braking. This is most easily done by turning the flight path downward so that the vehicle intersects the earth's atmosphere. The performance requirements for this maneuver are illustrated in figure 3. On this figure, the velocity increment required for immediate reentry is plotted on the vertical scale, with the flight time in seconds plotted on the horizontal scale. There are three trajectory cases considered on this figure.
Case 1 is for an optimum launch trajectory as given to us by the Marshall Space Flight Center. In this case, burnout velocity is achieved at an angle of +4° and the superorbital escape regime is entered at satellite velocity at a flight path angle of zero degrees. It can be seen that a great deal of velocity change may be required in order to effect an immediate reentry as escape velocity is approached. In order to improve this situation, we have considered alternate trajectories in which the initial path angle has been reduced.
In Case 2, the initial path angle of 4° changes to zero degrees in going between satellite and escape velocity. [18] In Case 3, the initial path angle of 6° at satellite velocity changes to 2° at escape velocity. It can be seen that for these two cases the velocity increment required to effect an immediate reentry greatly reduced. The amount of performance penalties associated with these alternate trajectories has not yet been determined.
It should also be pointed out that there may be another undesirable effect associated with these alternate trajectories. This has to do with the escape situation that would exist prior to the attaining of orbital velocity. With these alternate trajectories, the reentry angle following an abort at suborbital velocity will be much greater and, therefore, the reentry decelerations will, of course, be greater. How well the auxiliary propulsion system can cope with this situation is also yet to be determined. Thus, it is obvious that a great deal of study and analysis must be done in order to arrive at an optimum launch trajectory, taking into consideration the various abort maneuvers that may be required, as well as the performance of the launch propulsion.
Figure 4 outlines some of the aerodynamic considerations associated with the guidelines under discussion. As pointed out previously, the heating and loads criteria will undoubtedly be generated by consideration of the various abort and emergency maneuvers rather than the return from the moon in the normal mission. From a heating standpoint, it might be pointed out that the vehicle must be capable of reentries at escape velocity. Not only must convective heat transfer analysis be extended to this new velocity, but also a great deal must be learned about the radiation heating associated with the bow shock wave. There are indications that this heating may be on a level equal to that of the convective heating. It might be mentioned that the effectiveness of the ablation heat shield will be greatly reduced when coping with radiation heating as compared to the normal situation when convective heating is the criteria.
The needed aerodynamic capabilities of the vehicle will be determined by studying the trade offs between aerodynamic L/D, guidance capabilities, and the usefulness of the maneuvering rockets in making final corrections just prior to reentry. Also of concern is the ability to maneuver at low speeds after the vehicle has come out of the blackout region during reentry when no radio communications are possible. In the postblackout period, it is expected that another fix will be available from ground radar, and corrections for drift of the inertial systems aboard the capsule will be made. These corrections may require further maneuvers in order to land in the designated area.
[19] The actual aerodynamic configuration will be Compromised between these and other requirements spelled out throughout the guidelines. Needless to say, weight will be of primary importance in these considerations.
Figure 5 outlines some of the landing considerations resulting from the guidelines presented at this time. Inasmuch as the landing may be made either by a parachute system or by a glider, the landing considerations will be taken up separately for these two systems. For the parachute system it appears that the continued development of reliable large parachutes must be pursued. For the Mercury capsule, we found that no suitably developed parachute was obtainable and a development program had to be carried out in order to provide a design having adequate reliability. The advanced vehicle will be approximately three times the size of Mercury. It is desirable that some maneuvering capability be provided during parachute descent. This will be valuable in such events as a pad abort where a local hazard, such as a blockhouse, could be avoided during landing. Impact attenuation was found to be necessary for the Mercury capsule and in this new vehicle it will also probably be needed. Such devices as impact bags, collapsing struts, and honeycomb material should be investigated in order to provide sufficient latitude in choice.
In the event a glide vehicle is chosen, it will be necessary that an amphibious landing gear be provided inasmuch as both ground and water landings are a requirement. In addition, the L/D and wing loading criteria, which will provide a satisfactory approach and landing maneuver, must be determined in the same manner as was used for the Dyna Soar and X-15 projects. However, in this case, 30 knot winds and very rough water must be considered.
Finally, for the emergency and postlanding considerations, we must consider the effect of residual heated surfaces on the vehicle which might start local fires in the event of landing in brush or tall grass. Following water landings, the prime requirement is adequate flotation and water stability. Suitable location devices and survival support for the crew during the postlanding period will be needed following both water and ground landings in remote locations.
Figure 6 shows the requirements for the auxiliary propulsion system. Auxiliary propulsion propellants can either be solid or liquid, or possibly, a combination of these two. Inasmuch as there is a great difference between these two types of propulsion systems, the design [20] criteria will be different. The auxiliary propulsion system must provide for a wide variety of escape and guidance maneuvers. It is also estimated that the number of maneuvers during a mission may be large. These requirements can be met with a solid propellant system by use of a large number of rockets of various sizes. For a liquid system, restart capability will be required. In addition, it may be expected that more than one thrust unit will be required in order to provide sufficient thrust variation.
Reliability is a very important requirement inasmuch as the auxiliary propulsion system is the crew's "ticket'' back home. In the case of the liquid rocket system, it is not only important that the liquid fuel propulsion system meet the present design criteria, such as in the X-15 where the emphasis is put on safety (the rocket will not blow up), but it is also necessary that the rocket will start and function every time it is needed.
Since this is a major system, a certain amount of redundancy is justified. In the case of a liquid propellant system, extra thrust units and pumps will probably provide the minimum acceptable level of redundancy. For a solid propellant system, a suitable number of spare units will be required. In either case, a failure analysis is needed to provide the redundancy criteria.
