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TABLE I.-Candidate Experiments for Eros Rendezvous/Sample-Return Missions
Generic Mission phase of operation
mission. These instruments are only suggested as typical of what a rendezvous/ sample-return payload might be and are not a specifically recommended set of experiments. A description of another payload complement is found in the Meissinger and Greenstadt paper." A mass of 150 kg, including the 90 kg of Science and support equipment, is assigned to the trans-Eros and asteroid surface operations. This mass is jettisoned upon departure from Eros.
It is assumed that a number of stationkeeping maneuvers would be required following rendezvous with Eros before landing (or docking) on the asteroid. These maneuvers would achieve the necessary reconnaissance needed for landing site selection (preferably on one of the poles) and most importantly provide ample time for the remote sensing measurements indicated in table I at various locations around the asteroid.
The recommended sequence of maneuvers is illustrated in figure 1. They start at point 1, which is 1000 km from the asteroid in the solar direction. This position seems most compatible with rendezvous because the asteroid should be well lighted in a dark sky for optical tracking during final rendezvous. The sequence of maneuvers allows the spacecraft to approach the asteroid to within 250 km from three directions: Sun side, terminator, and dark side. The first two approaches should permit an accurate determination of the polar direction” (for landing) and provide good surface appearance data under limiting illumination conditions. The third approach would permit a search
Figure 1.-Stationkeeping profile for 1977 Eros sample-return mission. Dashed line indicates out-of-plane motion.
*See p. 543.
within a few degrees; see p. 133.
from the dark side for a dust cloud or atmospheric halo around Eros, using photometers. The final approach is recommended along the rotation axis to landing. The entire sequence would take about 55 days as shown in the figure. Remote teleoperator landing of a spacecraft is described by Meissinger and Greenstadt;” however, because of the large communication distance, response times are on the order of a half hour. The stationkeeping and landing operations were regarded as essentially the same for both the ballistic and low-thrust flight modes. The weight breakdowns for the baseline mission assume that the entire spacecraft (less outbound propulsion) is docked with the asteroid during sample collection and surface experimentation. A budget of 75 m/s has been allowed for all stationkeeping, docking, and separation maneuvers at Eros. A conservative estimate of 350kg is assigned to the landing mechanism and maneuver/docking propulsion inert mass. This mass is jettisoned upon departure from Eros.
PROPULSIVE ENERGY REQUIREMENTS
The analysis of round-trip interplanetary space missions generally requires the survey of compatible outbound and return trajectories. Whereas one-way orbiter or flyby missions may utilize near-optimum outbound trajectories, round-trip missions compromise the performance on both outbound and return legs so that the overall mission energy requirements are minimized. An expedient method of examining large quantities of round-trip trajectory
Figure 2.—Solar electric contour map for 1977 Eros sample-return mission. v1 is the hyperbolic excess velocity at Earth departure, v2 is the hyperbolic excess velocity at Eros arrival, and JVT is a measure of electric-propulsion energy requirements.
Figure 3.—Ballistic impulse contour map for 1977 Eros sample-return mission.
characteristics is shown in figures 2 and 3 for the 1977 opportunity to Eros as accomplished by an electric-propulsion flight mode and a chemical-propulsion flight mode, respectively.
ELECTRIC-PROPULSION FLIGHT MODE
A measure of the propulsive energy requirements for electrical rocket systems that employ finite-thrust trajectories is given by the time-integrated effect of the continually applied low acceleration. In figure 2, this energy measure has the symbol J with units of square meter per cubic second. The higher the J, the more difficult the mission. Mascy, Dugan, and Pitts (1968) have shown that the systematic mapping of this energy parameter provides a convenient technique for determining the best launch and arrival dates and the effect of varying trip times and stay times. For both figures 2 and 3, the launch and return dates at Earth are given along the abscissa and the arrival and departure dates at Eros are given along the ordinate. An example of an 1100 day mission is shown in figure 2 departing from Earth on Julian date 2443150, arriving at Eros 2443625, staying 100 days, departing from Eros 2443725, and returning to Earth on 2444250. In a similar manner, one can fashion missions of varying outbound or return-leg trip times, launch dates, stay times, etc.
The electric-propulsion system assumed in this report utilizes solar photovoltaic cells to convert the Sun's energy into electrical energy for acceleration of ionized mercury propellant. A power level of 10 kW and a thrustor specific impulse of 3000 s are used in the analysis. Similar to the systems and technology described in the TRW, Inc., report, the overall electric-propulsion module has a mass of 300 kg. An additional mass of approximately 400 kg has been assigned to the interplanetary bus, which comprises the engineering subsystems such as communications and data handling, central computer and mission sequencer, thermal and attitude control, and support structures. The launch vehicle used to inject the solar electrically propelled interplanetary vehicle onto a trans-Eros trajectory is the Titan IIID/Burner II.
CHEMICAL-PROPULSION FLIGHT MODE
In figure 3 is presented the contours of impulsive velocity increment, which is a measure of the propulsive energy requirements for chemical rocket systems that employ ballistic trajectories. This energy measure has the symbol Ay with the units kilometers per second. The higher the Av, the more difficult the mission. The outbound transfer contours in the lower left give the total impulse for two maneuvers: (1) impulsive escape from a 185 km (100 n. mi.) parking orbit at Earth and (2) impulsive rendezvous at Eros. In the background, dotted contours of declination of the launch asymptote (DLA) are given. Notice that the region of minimum total outbound impulse, Avo • 7.5 km/s, lies over absolute values of DLA that are greater than –70°. This implies a serious problem for launches from the Eastern Test Range because range safety constraints require that DLA 336°. Return transfer contours of impulsive departure velocity at Eros are presented in the upper right of figure 3. Dotted curves of Earth reentry speed of the return-sample capsule are shown in the background. The region of minimum departure impulse, AvR < 2 km/s, lies over Earth reentry speeds of less than 13.7 km's (45 000 ft/s), which should not pose a reentry technology problem. A sample mission that requires a total trip time of 3 yr is laid out with the arrowed line. The stay time at Eros would be almost 1 yr, 328 days. To use outbound and return transfers that are reasonably close to minimum required energies, it can ||