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returned to the landing vehicle, with round-trip communication delays of 35 to 40 min. A fully autonomous vehicle that would perform the final approach and landing at Eros without assistance by ground control would be more complex, more costly, and less reliable.

This paper discusses the scientific objectives to be achieved by an Eros landing and sample-return mission, the instrument payload to be carried, and the mission profile and critical mission phases to be executed. The conceptual design of a solar electric-propulsion spacecraft bus, or stage, capable of returning to Earth a capsule with 100 kg of Eros surface sample material will be described. The round-trip mission time is about 1000 days. The results of this investigation indicate the feasibility of this mission based on available electric-propulsion technology plus existing spacecraft design concepts and flight hardware. Such a bus could be developed in time to meet the 1977 launch opportunity. Similar opportunities occur approximately every other year.


It is believed that detailed examination of matter from an asteroid will provide information valuable to an understanding of the processes of planetary formation and of solar system hetegony (Alfvén and Arrhenius, 1970b). No foreseeable development of onsite techniques can hope to achieve the detail or depth of analysis that is possible with full laboratory equipment, as experience with lunar samples from Apollo has readily demonstrated. The central aim of an asteroid sample-return mission, therefore, will be to acquire asteroidal material to permit such an examination by laboratory analysis. It does not follow, however, that ambient or onsite measurements will have no place on an asteroid return-sample payload. Such measurements will contribute significantly to interpretation of laboratory results. Hence, instrumental surveillance of the target asteroid will be an important phase of the mission before, during, and after touchdown and perhaps after takeoff as well.

Surveillance will serve two functions, both of which encompass extensive local measurements. These functions are (1) selection of an optimal landing spot and (2) characterization of the physical context in which the asteroid is found and from which the samples are taken. These functions are naturally interrelated.

Several factors may enter the selection of an exact landing point. One major consideration involves the relative motion of the spacecraft with respect to the terrain at touchdown, which dictates landing near a pole of the asteroid's axis of rotation. Another consideration will be the angle of solar illumination. These are described in a later section. In addition, unpredictable properties of the asteroid may contribute to site selection. Among them would be local topography, which might determine that one area would be more level or more varied in composition than another or that samples would be obtained with more facility there than elsewhere, and local magnetic signature, which might suggest that samples from one spot would be more revealing of the early planetary environment than those taken from another.

The physical significance of laboratory analyses of returned samples could be seriously compromised without reasonably complete specification of the immediate and general present environment from which the samples come. It will be important to know how representative of the body composition and of the site materials the returned samples will be. For example, the cumulative effect of solar-wind impact on some surface materials may be influenced by a strong local magnetic field. Therefore, instruments measuring directly the local field, its gradient, and the resulting solar-wind deflection would be very important as part of the scientific instrument package. Complete field, particle, and optical characterization of the solar wind around the asteroid, of the asteroid as a whole, and of the landing site will therefore be essential for successful completion of the mission objectives.

To summarize, the overall mission purpose of collecting samples of asteroidal material from which comprehensive inferences on solar system formation can be obtained with minimal ambiguity will be served by three interrelated mission objectives:

(1) Examination of the asteroid's geometrical configuration and of its environment, including its interaction with the solar wind, if any

(2) Acquisition of samples of asteroidal material from the surface for return to Earth

(3) Observation and characterization of the site from which samples are taken and documentation of the relationship of the samples to the site and of the site to the asteroid

The scientific objectives described above will be served by groups of instruments that provide the following functions:

(1) Measurement of the ambient solar wind, the distant electromagnetic
properties of the body, and the interaction, if any, of the body with
the solar wind
(2) Observation of the asteroid's size, configuration, surface features,
rotation, and optical properties
(3) Detection of gaseous ionized envelope or plasma sheath
(4) Examination of the surface and subsurface characteristics at the
landing site
(5) Observation of surface features at the landing site
(6) Observation of ambient conditions at the site
(7) Acquisition of sample material

The instruments needed to perform the tasks included in the above functional categories are given in the following list. Asterisks denote items that might be assigned a lower priority than the others because their data would be unessential to success of the mission, being partially redundant in relation to what laboratory analysis would discover or being partially deducible from other sources.

(1) Ambient background and interaction measurements:
(a) Plasma probe
(b) dc magnetometer and gradiometer
(c) ac magnetometer
(d) Plasma wave detector
(e) Dust (micrometeoroid) detector
(f) Cosmic-ray telescope”
(g) Gravity gradiometer
(2) Asteroid observation:
(a) Imaging telescope (TV)
(b) IR, UV, and visible spectrophotometers
(c) Photopolarimeter
(3) Gas envelope detection:
(a) Low-energy plasma analyzer
(b) Ion mass spectrometer*
(4) Surface examination:
(a) Surface scraper
(b) Seismic detector, possibly with “thumper”
(c) o-Scattering analyzer"
(5) Surface observation: Imaging telescope (TV), as in function (2)
(6) On-site ambient environment detection: same as function (1)
(7) Sample acquisition:
(a) Loose matter collector
(b) Core borer

The scientific value of the mission would be enhanced if certain instruments were left on the surface together with the communication system required for telemetering their data to Earth. The weight of the devices left behind would be taken up, in part, by the collected samples. Detached instruments would include those for functions (1) and (5) plus the low-energy plasma analyzer (function (3)) and the seismic detector (function (4)). However, extended autonomous operation of a telemetry system and its power source on the asteroid with communication distance to Earth in excess of 2 AU involves technical problems not considered within the scope of this paper.


Trajectories, performance characteristics, and payload capabilities for one-way and round-trip missions to Eros have been investigated by Friedlander, Mascy, Niehoff, and others (Friedlander and Vickers, 1964; IIT Research Institute, 1964; Mascy and Niehoff?) for both ballistic and low-thrust

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propelled vehicles. Figure 1 shows representative outbound and inbound trajectories of a 1050 day round-trip mission with a 50 day stopover at Eros for the 1977 launch opportunity, based on data obtained by Mascy.” The mission uses solar electric propulsion both ways, with thrust characteristics and thrust pointing angles optimized to return a maximum amount of asteroid sample material to Earth. The vehicle is launched by a Titan IIID/Burner II booster and uses 10 kW of initial propulsive power at Earth departure. Low thrust is applied continuously during the outbound phase such that the vehicle arrives at Eros with zero relative velocity voc and can land on the asteroid with almost no additional propulsive effort. Similarly, low thrust is applied continuously during the return trip to reduce the approach velocity on returning to Earth and the required Earth capture maneuver. We assume that the sample-return capsule carried by the interplanetary bus vehicle will be inserted into an eccentric Earth parking orbit for subsequent retrieval by orbital shuttle or by a deorbit maneuver, atmospheric entry, and parachute landing. This mission profile is shown schematically in figure 2 and is used as a basis for defining the vehicle design features and operational characteristics to be discussed below. We note in figure 1 that the outbound trajectory departing from Earth on February 25, 1977, swings in a wide arc to an aphelion distance of 1.67 AU to achieve the desired velocity matching with the target at the encounter date of July 10, 1978, near perihelion. A gradual plane change necessary to attain the 10:8 orbital inclination of Eros is included in the outbound propulsion phase.


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Figure 1.-Eros round-trip trajectory.

*Specific data used in this article are based on work by Mascy and Niehoff and are essentially in agreement with data published in their paper in this volume on p. 513.

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Figure 2. —Eros round-trip mission profile (schematic).

The return trip departing Eros on August 29, 1978, and arriving at Earth on January 12, 1980, has similar characteristics. Mission opportunities with comparable characteristics occur about every 2 yr. 1977, 1979, and 1981 are favorable mission years (Mascy and Niehoff"). A characteristic feature of this class of mission profiles is the large communication range to Earth (2.1 AU) and the fact that Earth and Eros are in almost exact opposition at encounter. These conditions do not change much during the 50 day stopover because Earth and Eros move nearly at the same rate. In the reference trajectory, the arrival at Eros occurs a short time after syzygy. Communication blackout must be avoided during this critical part of the mission. The Earth-Sun separation angle subtended at Eros is initially 3°. This gives a margin of only 1° from the blackout zone, 2° on both sides of the solar disk, which is assumed under conditions of average solar activity. Actually, because during the late 1970's solar activity will be increasing toward a maximum level, a larger margin than 1° would be desirable. The separation angle increases to 5°5 during the 50 day stopover. Therefore, a 20-day delay in arrival will increase the margin by 1°. This can be achieved with only a minor change in payload performance owing to the flexibility of low-thrust missions, as shown by Mascy.” A delay in arrival date is also desirable to improve seasonal conditions at the preferred polar site as discussed below.


Several major performance advantages accrue in this mission from the use of electric propulsion. The first, and by far the most important one, is the large

*See p. 522. *See p. 525.

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