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Electric thrust is sufficient to hold the vehicle in stationary position against the small local gravity during hover phases 1 and 2. A set of small hydrazine rockets provides acceleration and deceleration on descent from hover position 2 to position 3. At hover position 3, located 200 m above the terrain, a final check of the landing site is made to take corrective action for obstacle avoidance, if necessary. This 40 min hover phase requires use of the hydrazine rockets in an intermittent thrust mode. The final descent to touchdown ends in a free fall from about 50 m altitude, resulting in an impact velocity of about 0.7 m/s based on an assumed gravity acceleration of 0.5 cm/s”. Autonomous control of the descent phase velocity profile from an altitude of about 10 km is achieved by means of a radar altimeter and a three-beam Doppler radar system that operates in a manner similar to the Apollo lunar module landing radar. The threshold velocity detectable by this system is 1.5 m/s. The descent is in nearly vertical direction, which simplifies the onboard computation of velocity and attitude corrections. In addition to the radar system, the vehicle uses a vertical attitude gyro as a redundant attitude control reference. The solar array is designed for retraction by the rollup storage mechanism. After completion of the final electric thrust operation, about 2 hr prior to landing, the solar array is retracted for protection against dynamic loads at impact but with a sufficient portion of array paddles protruding to provide about 200 W to operate housekeeping systems and the landing radar. The liftoff and return flight sequence will not be discussed in detail. In principle, this sequence is a reversal of the outbound transfer but with the approach guidance to Earth made simpler by the absence of target ephemeris uncertainty and by the availability of a DSIF station on the ground.
VEHICLE CONFIGURATION AND DESIGN CHARACTERISTICS
Conceptual configurations of the solar electric spacecraft during cruise and after landing are shown in figure 7. The vehicle consists of a center structure that houses the electric-propulsion module, engineering subsystems, scientific instruments, sample collection tools, and the sample-return capsule. Attached to the center body are two pairs of lightweight solar array paddles that are deployed from storage drums in window-shade fashion by means of extendable tubular booms. The flexible landing gear consists of four legs having footpads lined with crushable material for absorbing impact energy as in the Surveyor spacecraft. Spring-released anchoring devices, not shown in the sketch, are used to secure the vehicle's position after touchdown under the extremely small surface gravity of Eros. This configuration is derived from an earlier solar electric spacecraft design study (TRW, Inc., 1970).
The electric-propulsion module, mounted opposite the payload bay, is shown in greater detail in figure 8. It consists of electric power conditioning units (PCU’s), an array of electron-bombardment mercury ion thrustors
mounted on a flexible support fixture for thrust vector control, a propellant tank, and a feed system. The high density of the mercury propellant permits storage of a 750 kg propellant load in a 0.5 m diameter compact spherical tank located near the vehicle's center of mass. This minimizes the change of mass distribution and its effect on attitude control during cruise and landing. Five ion thrustors are provided, only two of which are normally in use during the outbound and inbound transfer phases. The three remaining thrustors serve as standby units to assure a high system reliability in a mission that requires a total thrust time of 1000 days. The delicate thrustor system uses a shroud for protection against dust stirred up during touchdown and surface operations. Dust covers must also be provided for the apertures of optical and other SensOrs. The rollout solar array with each paddle measuring 2 by 14 m when fully extended provides 12 kW of initial power at 1 AU, 10 kW of which is used to operate the electric-propulsion system. The remaining power is used for housekeeping and telemetry and includes a 10 percent margin against contingencies such as solar array performance degradation due to solar flares. Prior to landing, the paddles are retracted for protection against the landing impact and flying dust. Subsequently, small paddle segments are extended to generate power of about 400 W for surface operations, housekeeping, and high-data-rate telemetry of TV pictures. After takeoff from Eros, the array is again fully extended for the return cruise. As shown in the design illustration, the four solar paddles can be rotated around their deployment booms to improve array illumination primarily during the approach, descent, and hover phases, and after landing. During the transfer phase, small changes in solar array orientation relative to the center body are useful but not required, permitting an additional degree of freedom for optimum thrust vector pointing. During the landed phase, array reorientation may be required to accommodate changes in Sun elevation. By splitting the array into four narrow paddles instead of two, field-of-view obscuration of the optical sensors and the high-gain antenna due to paddle reorientation can be avoided. These deployment and orientation sequences are compatible with design specifications of the rollout array shown in the design illustrated. A full-scale 2.5 kW engineering model of this array design has been developed for NASA by General Electric and has successfully completed extensive electrical, thermal, and mechanical tests in the laboratory. A set of four differentially throttlable hydrazine thrustors, each with a maximum thrust level of 4 N (1 lb), are centrally mounted on the underside of the center body. They provide thrust required during the descent and ascent phase and support the vehicle during the extended final hover phase prior to touchdown. With an assumed small surface gravity of 0.5 cm/s2, the total hydrazine propellant consumption for a 40 min hover period is 16 kg for a vehicle of 1400 kg gross mass. The total maneuver sequence performed by the hydrazine thrustors requires about 40 kg of propellant, equivalent to 60 m/s of Av expenditure. A variety of optical sensors are required to control vehicle operations on approach to the asteroid, during the final descent phase, during landed operations, and during ascent. The use of a TV system is required both for the vehicle control operations and the scientific observations that follow:
(3) Obstacle avoidance by command control from Earth
These operations will require wide-angle and narrow-angle, high-resolution TV image systems. Three or more separate TV systems mounted on two-axis gimbal platforms with different and overlapping fields of view are envisioned. For landing site inspection and selection an image system with at least 0.0014° resolution will be required comparable to the planetary image system of the TOPS spacecraft. The attitude reference sensors include fixed, fine and coarse, Sun sensors, and a one-axis gimbaled star sensor that uses an electronic image scanning principle similar to the Mariner star tracker. This device permits use of alternate reference stars that come into view at different times in the mission and can be tracked without interference by the long solar array paddles and other deployed appendages (Meissinger and Benson, 1970). An additional optical sensor is required for locking the high-resolution TV camera on the asteroid during the approach and hover phases when successive TV frames of the polar area must be taken for determining the location of the pole and selecting a landing site. The communication system operates on S-band and uses design principles developed for and successfully used on Pioneer spacecraft. The 2.4 m diameter parabolic high-gain antenna dish is mounted on a hinged deployment arm that permits Earth pointing in all directions relative to the vehicle body without obstruction by other deployed appendages, using a two-axis rotation joint. The same deployed configuration is used during cruise and landed operations. In addition to the high-gain antenna, a pair of omniantennas is provided, one on each side of the center body, to maintain an uplink command capability in all vehicle attitudes. A 100 W solid-state transmitter composed of four parallel 25 W channels provides incremental power options desirable for flexible use of the telemetry system. Low telemetry power and bit rates are used at times when power is needed primarily for propulsion purposes. High bit rates are available for telemetry of TV images during critical mission phases. Table I lists representative bit rates and communication intervals per TV frame at 2.1 AU communication range for 25 and 100 W of transmitter power with ground coverage by 25.9 and 64.0 m DSIF antennas. The effect of TV image data compression by a 2:1 ratio is also shown. The unprocessed TV image is assumed to contain 2.5 × 10% bits. Bit rates of 4096 bps are available for telemetry to the 25.9 m ground station and 65536 bps to the 64.0 m ground station using 100 W of transmitter power. This corresponds to telemetry intervals of 610 and 38 s per TV frame without data compression and to 305 and 19 s with data compression, respectively. The lower time intervals are quite small compared to the 17.5 min one-way transmission time delay from Eros. A principal advantage of the electric spacecraft is the ability to meet high data
TABLE I.—Eros-to-Earth Communication Data Rates
Downlink Single TV frame
Transmitter DSIF Unprocessed With 2:1 data power, W antenna, m data compression 25.9 1 024 2440 1220
25.9 4.096 610 305
64.0 16384 152 76
64.0 65536 38 19
Average communication range during stopover: 2.1 AU; S-band telemetry; 2.4 m spacecraft antenna; TV frame contains 2.5 x 10° bits (unprocessed).
rate requirements without demanding coverage by the 64.0 m DSIF station, owing to the large unused power capacity available for telemetry during critical mission events. The following list is a summary of system characteristics based on performance data from Mascy and Niehoff.”
(1) Launch vehicle: Titan IIID/Burner II
Mass estimates of the electric bus vehicle and the sample-return capsule, also based on the performance data of Mascy and Niehoff, are given in table II. By holding the return sample mass to 100 kg, a 10 kW bus vehicle launched by Titan IIID/Burner II has adequate performance margin. A much larger sample mass of 200 to 300 kg can be returned by using a higher powered bus vehicle (15 kW), which would require the more costly Titan IIID/Centaur booster. We conclude this discussion with a chart (fig. 9) showing the current status of critical subsystems that are required to develop the electric bus vehicle for this mission. In all categories except the solar array, a first generation subsystem with adequate performance to achieve the mission has been flight proven and would be ready for application to the system. Technology
7See p. 524.