circular in cross section, but with a diagonal, and therefore elliptical, base. The rearward cone, shown pcinting upward in the figure, is elliptical in cross section, but has a base normal to its axis. The arrangement provides a high lift/drag ratio, at a constant, zero angle of attack. The lift can be rotated 360° around the flight axis by rotating the capsule. The high lift/drag increases the depth of the entry corridor, improving the margin for error, and permitting high entry speeds. It also reduces the peak acceleration, and the peak heating load. It provides a wide selection of landing places, regardless of time and direction of entry. This would be most advantageous in returning from a planet, permitting the crew to maneuver for minimum propellant consumption and maximum accuracy, rather than for exact timing of arrival at earth. K It is proposed to use the large tonnage of liquid metabolic waste for boundary layer cooling during reentry. The aerodynami.c pressures on the nose of the capsule arc small compared with precsures in rocket combustion chambers, so that a pressure feed system using a low-pressure tank would be suitable. Indeed, the pressure is so low that variations in back pressure across the injection area can be overcome by simply providing an excess of pressure drop for the worst case. The orifice sizes could then be adjusted permanently for all flight conditions, since the angle of attack remains constant. The capsule could be tested during the practice flights in which the launch rocket is recovered, as provided in Section 4. (It would also be necessary to flight test the APOLLO capsule, if it were used, to qualify it for the higher speeds it would encounter in returning from interplanetary expeditions). It is not contemplated that artificial gravity will be needed during long space trips. However, if gravity is found necessary, the procedure would be to launch the spacecraft in pairs, and after they are on heliocentric orbit, where rendezvous is simple, to bring them together, base to base, attach them to one another rigidly, and then upin the biconic structure end over end by means of small jets. In most cases, the second ship could be a freighter ship already scheduled in Section 4 to make the trip anyway. 22 2.3.3 MAINED LANDING These are itemized under two headings in Section 4, namely for manned landings at semi-permanent bases on the moon and on Mars. Four landings per year are arbitrarily allotted to the moon, enough for a landing every 3 months at a single base, or every 6 months at each of two bases. The lunar landers could be launched without the use of tanker support. They would carry men directly from the earth to the lunar surface, and back to the earth. Three landings per synodic period of 2.13 years are allotted to Mars. Two would be sent on conjunction-class trips, and one on another class of trip, typically Venus swingby. One crew would complete the standard conjunction-class trip. Another would begin on a standard conjunction-class trip, but return on the return leg of the next short round trip. Meanwhile, the outbound crew on the short round trip would return on the next conjunction-class return leg. In this way the base could be kept continuously manned, with overlapping st ys, without the need for any crew to remain for two conjunction-class trip times. The simple conjunction-class trip would be auxiliary to the others, both for insurance and expanded capability. The Mars landing trips would differ from the lunar landing trips in the number of launches required for each crew. Each passenger ship would require two launches, one being a tanker. The lander would be a separate freighter, requiring no tanker. Each crew would therefore require three launches, but only two space craft. The separate freighter would rendezvous with the passenger ship on a Mars parking orbit, and provide the means of landing and return to the passenger ship, which would bring the crew back to earth. The entire lunar lander would return to earth, where an aerodynamic entry capsule would be deployed for entry into the earth's atmosphere. Only a part of the Mars landing craft would return to the other ship in the Mars parking orbit, namely an ascent rocket carried along for the purpose. The earth reentry capsule would be carried in the capture-only spacecraft. In spite of these differences, the lunar and Kars landers would use the same structures, Instead of sending the nose cone into a near-parabolic orbit one could send it into a low-altitude, circular parking orbit. There would be enough propellant for a retrothrust landing, or the standard aerodynamic entry capsule could te landed. The vehicle would be useful in global transportation experiments, and at the same time could be used to perfect the aerodynamic entry vehicle for use in returning from space missions. Global transportation experiments could be combined with earth orbiting laboratory operations, and with training flights for astronauts and scientists. They could also be combined with the deployment of unmanned vehicles such as terrain scanners, etc. 2.4 MASS TABULATIONS A salient feature of the plan proposed in this paper is the use of a single launch rocket-space rocke design for all space missions. The propulsive stages of the space rocket shown in Figure 2.2-a are sized according to spacifications for a standard space ship mutually adjusted for circular-capture, mauned expeditions to Venus and Mars at each synodic period of each planet. The mass tabulations are shown in Tables 2.4-a, 2.4-b, 2.4-c, and 2.4-d. The mass fractions given in the specifications are for the worst cases which will ever need to be encountered in making a trip to the indicated planet. They are taken from the data presented in Section 3. The mass fraction shown for departure from the earth are for departure from a near-parabolic trajectory, rather than from a circular parking orbit as shown in Section 3. Table 2.4-e relates the mass tabulations from Tables 2.4-a to 2.4-d to other possible celections. None of the other sizes rival the 263 metric-ton size in economy or overall simplicity, for Venus and Mars exploration. The real criterion, however, is in whether the 263-ton size will also do the other misions we may want it to 24 TABLE 2.4-a · MASS TABULATION, MUTUALLY ADJUSTED FOR MARS AND VENUS SHIPS. Enter atm. at Earth, 3400 kg + 5600 kg water At a mixture ratio of 5.5, the first stage above contains 25 TABLE 2.4-b MASS TABULATION, MUTUALLY ADJUSTED FOR MARS AND VENUS SHIPS. Enter atm. at Earth, 3400 kg + 5600 kg water 9000 At a mixture ratio of 5.5, the first stage above contains |