During the course of this design study the overall configuration of the probe underwent a number of changes. As design work on the payload and subsystems progressed the overall requirements that the probe would have to meet became clearer and thus the constraints on its configuration more explicit. This report has necessarily concentrated on the final configuration of the probe but some discussion of how it was arrived at is in order.

Initial Design. The basic exploration of the probe's mission requirements made it clear that it would require a number of basic design features:

• A large-aperture optical assembly for imaging the target during flyby.
• A high-gain antenna for rapid download of large quantities of image data.
• Propulsive capability for trajectory adjustment.

At this stage a number of factors had yet to be decided, such as whether a separate trajectory injection stage would be used. However, it was already evident that the probe would be too small to carry the optical assembly required on a scan platform. Figure 14.1 shows an early candidate probe configuration, with a diameter of about 1.5 m.

At this stage power systems had not been considered in detail, so no solar arrays are shown. A relatively large on-board fuel supply is also shown as it was considered that the probe may provide much of the injection delta-V as well as trajectory correction.

Evolution of Design. Subsequent study concentrated on a design with a separate boost stage, reducing the size of the on-board fuel supply. Analysis of the communications link budget also showed that a high-gain antenna diameter 0.5 - 1.0 m would be acceptable. By now the encounter slew had also been modelled in more detail and it became apparent that although a scan platform was not feasible some form of scanning periscope may be preferable to slewing the entire probe. From this a revised configuration was produced, of which Fig 14.2 shows a side cross-section.

This configuration appeared promising, but on further examination had a number of disadvantages. The use of a rotating periscope mirror added a complex mechanism to the probe which would have to reliably survive launch acceleration. The choice of a high-thrust solid propellant boost stage on grounds of design simplicity made this a particularly awkward area. The use of a scanning periscope also effectively precluded the use of the front face of the probe for mounting an antenna. This meant that the antenna would have to either be side-mounted (limiting its size and causing centre of mass problems) or rear-mounted (interfering with the propulsion system). It was thus decided to use a fixed imaging sensor and track the target by slewing the entire spacecraft.

By this point the probe was approaching a final configuration: a short, wide cylinder (or box or prism) some 1.0 -1.5 m wide and 0.3 - 0.6 m deep, with the imaging sensor looking radially outwards and the high-gain antenna on the forward face. Thought now started to be given towards power supply and thermal control.

The nature of this probe's mission meant that solar cell power supply was potentially awkward. Most deep space missions are designed for a particular encounter geometry with respect to the target and the Sun. It is often therefore possible to use a fixed solar cell configuration that provides adequate solar power for this geometry. This mission though, by being designed to cope with a range of targets, could not rely on any such particular encounter geometry. By making the solar arrays deployable it was possible to increase the total projected area over a range of Sun angles . The configuration of Fig 14.3 shows one such design, which allowed for Sun angles from side-on to front-facing. Modelling of this configuration indicated that for solar arrays tilted 15° from the side faces the projected solar array area was always between 0.8 and 1.5 of that of a single array for any sun angle up to 120° away from the forward axis of the probe.

Nonetheless this configuration still required several array deployment mechanisms and could still not cope with a full range of Sun angles. An improved configuration that needed no deployment is shown at Fig 14.4. By having both a cylindrical array around the spacecraft body and an annular circular array surrounding the high-gain antenna it was possible to provide power for a wide range of Sun angles.

How large should the front array be as a fraction of the front face? For initial design purposes it was assumed that the transmitted downlink power would be a fixed fraction of the array power and thus proportional to array area. The antenna gain was proportional to its area, and so the transmitter EIRP was proportional to the product of the front array and antenna areas. This was a maximum when they were both equal, i.e. when the antenna diameter was 0.7 of the probe diameter. Although later design work concentrated on a fixed transmitter power this size ratio was maintained as a good balance between array and antenna size.

At this stage concerns regarding Sun impingement on the image sensor led to a constraint of Sun angle from forward axis of no greater than 60° being adopted for a time. On this basis the probe diameter/depth ratio for even power as a function of Sun angle was calculated as about 3:1. Based on rough power estimates a probe size of 1.5 m diameter and 0.5 m depth was thus obtained and initial thermal design was carried out (see Section 16). When modelling of asteroid encounters was carried out in detail though, it became clear that in many cases the probe would need to have a Sun-axis angle of about 90° to keep its antenna Earth-pointing during cruise. Although this was not an essential requirement it was viewed as highly desirable and so the probe was redesigned to allow this. At the same time it was decided to add an annular array to the rear face as well to allow operation at any Sun angle.

Final Probe Configuration. The final probe configuration is shown at Fig 14.5. As discussed in Section 16, the probe design was modified to be narrower and deeper to avoid excessive heat loss, and the side solar array had thermal control surfaces included. Note also that the aperture for the imaging system requires a cover to protect it from direct sunlight during the cruise phase of the mission. This cover would have to be jettisoned as the encounter approached; although this had the unwelcome effect of introducing a mission-critical mechanism into the probe it was felt that the benefit of removing constraints on pointing angles justified doing so.

This final configuration, with a probe diameter of 1 m and depth of 0.5 m ensured that the power provided by the solar arrays was never less than 73W, at specified in the Section 15). Thermal modelling also showed that it allowed the probe's internal temperature to remain in the range 270 - 315 K over the full range of operating environments.

With the probe configuration specified its structural design was considered in more detail (Section 17). From this the probe's internal structure could be specified as in Fig 14.6. The probe's systems have been laid out such that the most massive (i.e. imaging system optical assembly and fuel tanks) are located at the probe centre of mass or symmetrically around it. Note that this diagram is still somewhat schematic in that the exact locations of some systems (especially electronics) has not been specified. It is also worth noting that the configuration of the probe's imaging system requires a viewing aperture not only in the outer side wall but also in the thrust cylinder. One important aspect for a more detailed structural design would be to examine the implications of this for the thrust cylinder's strength and design.