European Solar Polar Orbiter Mission

Remaining Spacecraft Sub-Systems

The on-board memory requirements are defined in part by the maximum data latency setting of one week, thus the spacecraft acquires just over 2 Gbit of data between downlinks. Furthermore, it is required that the solar sail be highly autonomous, necessitating sizeable computational and storage capabilities.
Within the remaining sub-systems no significant technology issues were identified within this analysis. Note that the placement of the spacecraft within the plane of the sail film means that the heat generated by the sail, both reflected and emitted, has a very low view factor with respect to the spacecraft systems. Thus the sail thermal input to the spacecraft is negligible. An initial radiation analysis suggested a conservative design tolerance of 75 krad should be adopted.
The spacecraft configuration is assumed to be a cube of sides 1.1 m. We note however from the solar orbiter mission studies that the spacecraft thermal design may be aided by adoption of a rectangular design rather than a cube. Figure 2 shows a visualization of the SPO spacecraft in deployed configuration, allowing the volumetric requirements of the spacecraft to be analyzed and aiding in the launch vehicle configuration analysis later within this paper. The imaging science instruments are mounted internally, with field-of-view towards the Sun along the negative X-axis, this provides a clear field-of-view as the face is maintained in a sunward orientation through bias torque. A 2-axis steerable HGA is mounted on the anti-Sun side, the positive X-axis, so the HGA is provided with a degree of thermal protection due to the shadow from the main spacecraft body. However, in this initial configuration there is potential for the (hot) solar arrays to impinge the field-of-view of the HGA. Two sets of one degree of freedom steerable solar arrays are mounted along the Z-axis, either side of the main body of the spacecraft, allowing solar aspect angle to be varied during the cruise phase of the mission while the spacecraft is attached to the solar sail. The solar arrays are small, thus they are stowed against the ± Z-axis faces during launch. The negative X-axis, which faces the Sun during the science phase of the mission, is provided with additional thermal protection. However, all spacecraft faces are required to have sufficient thermal control, as during the cruise phase of the mission the sail attitude may be such as to expose any spacecraft surface to the Sun for a short period of time. The –X face mounts the spacecraft onto the solar sail. By mounting the spacecraft via the –X face we shield the science instruments from the deep space environment until after sail jettison, thus helping to maintain optical surfaces in optimal condition. This configuration however eliminates the potential use of these instruments during the cruise phase of the mission.

Figure 2 Solar Polar Orbiter preliminary deployed visualization

Variation of Minimum Solar Approach Radius

The spacecraft systems detailed previously were designed for a minimum solar close approach of 0.48 AU. However, from a trajectory perspective an optimal solution can be found by allowing closer solar approaches during transfer. As such, it is important to quantify the effect of varying the solar close approach radius on the spacecraft sub-systems, in order to define this sensitivity. We see in Figure 3 the effect of varying solar approach radius on the wet mass of the spacecraft, without the sail. The design points are intended to provide a good first approximation and as such are suitable for preliminary design analysis. The information in Figure 3 will be coupled with trajectory and sail design information later within this paper in order to fully quantify the effect of varying the minimum solar approach radius.

Figure 3 Variation of total spacecraft mass as solar approach thermal limit is varied

Required Sail Slew Rates

From trajectory analysis we note that inclination cranking constitutes the bulk of the transfer trajectory to the solar polar orbit. During the cranking phase the sail pitch is fixed at arctan(1/√2), while the sail clock angle flips from 0 deg to 180 deg.Error: Reference source not found However, it is clear that the sail thrust vector cannot be rotated through ~70.5 deg instantaneously. We thus investigate the effect of varying the sail slew rate in order to quantify requirements on the sail attitude control system. By examination of the locally optimal inclination control law¹¹ we anticipate that sail slew requirements should be rather low. The rate of change of inclination will approximate a signum function of a cosine curve, with the required sail slew maneuver naturally occurring when the rate of change of inclination is low.
Using a heliocentric trajectory model which includes orbit perturbations due to the terrestrial planets and which models the Sun as a finite uniformly bright disk and the sail as an 85 % efficient reflector, we propagate the orbit cranking sail trajectory for half an orbit revolution. We compare the inclination change over half an orbit against the instant sail slew scenario, allowing investigation of sail slew rates as seen in Figure 4. Figure 4 shows the drop-off, or degradation due to the finite sail slew rate on the rate of change of inclination. We see that at lower sail accelerations the degradation for a given slew rate is increased, while at lower solar radii the degradation is also increased due to the shortened orbit period. We see from Figure 4 that above a sail slew rate of 10 deg per day (10^-4 deg s^-1) the degradation of inclination change is less than 0.5 % at all the accelerations and solar radii considered. We thus define the required sail slew rate as 10 deg per day.

Figure 4 Rate of change of inclination with respect to instant slew scenario. Orbit radius indicated top-left, for three characteristic accelerations at each radii, 0.4 mm s^-2 (-), 0.5 mm s^-2 (--) and 0.6 mm s^-2 (···).