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not have a significant effect on payload over the range selected. Rather the payloads fluctuate a great deal because of the varying relationships between relative orbit nodes and perihelion from one launch opportunity to the next. It is clear that missions to any of these five asteroids (and of course many others) are possible with solar electrically propelled spacecraft launched by a Titan IIID/Centaur or an Atlas/Centaur and that a substantial scientific payload can be delivered.

Eros Trajectory

To determine requirements on the design of an asteroid rendezvous spacecraft, some details of a flight to Eros have been generated. The selected launch date” is in December 1974 and the launch vehicle postulated is the Atlas/Centaur. The spacecraft is equipped with the solar panels developing 10 kW at 1 AU and, except for a 10 day coast period in the middle of the flight, the engines thrust continuously for the 360 day trip to Eros. This combination of launch vehicle, power, and flight time allow a net spacecraft mass at the target in excess of 500 kg. This includes 50 to 100 kg of science instruments. A breakdown of the spacecraft mass at launch is given in table III. Figure 4 shows the relative positions of Earth, the spacecraft, and Eros projected into the ecliptic for various dates along the trajectory. After May 1975, the spacecraft and Eros are in the same position to the scale of this drawing, but a good deal of their separation is normal to the ecliptic and hence does not show in this projection. The most vital part of the flight from the standpoint of scientific return begins when the spacecraft approaches the asteroid. In this phase there are a number of problem areas, two of which are discussed below. To begin with, the position of the asteroid is not sufficiently well known for spacecraft navigation purposes. We expect an a priori error on the order of a few thousand kilometers. Because this is much larger than the intended closest approach distance to Eros (which is on the order of the dimensions of Eros

TABLE III.—Spacecraft Mass Budget

Component Mass,
Power and propulsion system (10kW) 300
Mercury propellant 150
Net spacecraft (including scientific payload) 500
Gross spacecraft mass at launch 950

*There are no firm NASA plans to launch an Eros mission at this date; it is used for example purposes only.

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Figure 4.—Eros rendezvous ecliptic plane projection. Thrust is fixed at 90°, o = 10 kW. Is = 3500, and to - 360 days. Launch date is December 7, 1974. The launch vehicle is the Atlas/Centaur.

itself), the error must be corrected. Before a flight, the ephemeris of the target would be improved to the largest extent possible by ground-based observa. tions.” Further improvement can be made by correcting the ephemeris of the target by using optical observations taken from the spacecraft while in flight. This procedure was tested on Mariners 6 and 7 and it is scheduled to be operationally demonstrated on the Mariner-Mars 1971 orbiter arriving late this year. It is a vitally necessary component of the Outer Planet Grand Tour Flight. Using typical optical system performance, it should be possible to begin resolving Eros when the spacecraft is 200 000 km away although it may be visible as a bright object well before this. The position of the asteroid against the star background, when combined with Earth-based radio-tracking data from the spacecraft, provides sufficient information to determine the space. craft's position relative to the asteroid. The information can then be used to reprogram the thrust history for the remainder of the flight, thereby correcting any errors. Another problem that arises is due to the fact that the thrust beam must be directed nearly toward the asteroid near the end of the flight. This can be seen from figure 5, which shows the asteroid-spacecraft-thrust beam angle as a function of date throughout the mission. Note that the angle goes to zero as the spacecraft approaches the asteroid. (This phenomenon is characteristic of any rendezvous trajectory because the spacecraft must slow down with respect to the target on the approach.) Thus observations from the spacecraft during

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this phase would have to be made through the exhaust. Two possible solutions to this problem are (1) to momentarily cease thrusting on the approach and reorient the spacecraft to allow observations or (2) to initially rendezvous with a point some distance from the asteroid, take the appropriate observations, then slowly close in on it. These and other methods are currently under study.

ACKNOWLEDGMENTS

The authors are indebted to many members of the staff of the Jet Propulsion Laboratory for their contributions to the work on which this article is based. In particular, we would like to thank J. M. Driver, C. G. Sauer, and Dr. C. L. Yen for their assistance in generating the mission analysis data. This paper presents the results of one phase of research carried out at the Jet Propulsion Laboratory, California Institute of Technology, under contract NAS 7-100, sponsored by NASA.

REFERENCES

Bartz, D. R., and Horsewood, J. L. 1969, Characteristics, Capabilities and Costs of Solar Electric Spacecraft for Planetary Missions. J. Spacecr. Rockets 6(12), 1379-1390.

Han, D. W., Johnson, F. I., and Itzen, B. F. 1969, Chebychev Trajectory Optimization Program. D2-121308-1 Final Rept. (NAS2-5185), The Boeing Co.

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SAMPLE-RETURN MISSIONS TO THE ASTEROID EROS

ALFRED C MASCY
WASA Ames Research Center
and
JOHN NIEHOFF
//T Research Institute

MISSION SELECTION

Of the many asteroids that are cataloged, selection of a target planetoid was narrowed to the Mars-crossing asteroids. The main belt and Trojan asteroids are much farther away and require longer voyage times and propulsive energy to accomplish sample returns. Within the Mars-crossing group, those asteroids were surveyed that have orbits inclined less than 15°, perihelia within 0.2 AU of Earth's orbit (and therefore require less propulsive energy), and diameters greater than 1 km (to assist terminal observation and rendezvous). Eros was chosen as representative of this class.

The mission consists of the following phases: (1) Earth departure; (2) trans-Eros trajectory; (3) Eros approach, rendezvous, site selection, topographical coverage, and docking; (4) surface operations including sample acquisition; (5) Eros departure; (6) trans-Earth trajectory; (7) Earth approach, capsule separation, and orbit capture maneuver; and (8) in-orbit quarantine until deboost command or orbit retrieval.

CANDIDATE EXPERIMENTS

A number of scientific instruments were considered in formulating a baseline Eros rendezvous/sample-return mission. These candidate instruments are listed in table I along with related science measurables, the mission phases during which they would operate, and an indication of what their mass and power requirements are expected to be. The instruments are divided into two categories: (1) those instruments essential for sample return and (2) those desirable for a viable rendezvous and landing science mission. Because not all instruments would operate simultaneously, the total power requirement is somewhat less than the accumulated one. By far the most important phases of the mission for instrument operation are stationkeeping and landing. Because the total science payload is only 90 kg, it was assumed that all instruments presented in table I would be included in the science payload of the baseline

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