Power for space probe missions has traditionally been provided in one of 3 main ways; battery, photovoltaic (solar) cell, and radioisotope thermal generator (RTG). Other systems, such as fuel cells, have been used to provide power in space, but not on such missions. The main features of these techniques are described below:

Battery Power. Batteries can be used as the sole source of power for missions combining low power requirements with short active lifetimes. Note that the lifetime constraint need apply only to the operational phase of the mission, as it can be preceded by a prolonged dormant phase with minimal power requirements. For example, the Galileo Jupiter entry probe was released 6 months prior to encounter, with the probe's batteries being used to power it only for the final few hours of entry and measurement.

Various sorts of battery technology have been used on space missions [Larson92]. Missions requiring secondary (rechargeable) cells have traditionally used nickel-cadmium batteries, although in recent years these have been supplanted by nickel-hydrogen cells on larger spacecraft. Where single-use batteries suffice silver-zinc cells have often been used but recent developments in lithium-based batteries (e.g. lithium thionyl chloride) have led to their use on a growing number of missions. Indeed, such batteries are now being considered for use in a rechargeable role [Dudley97]. Lithium-based batteries can now reliably offer power densities exceeding 200 Whr/kg.

Battery-based power systems are simple and compact, but can supply only limited amounts of power. Whilst such a system could probably power a spacecraft during a short flyby encounter it is unlikely to be able to support a prolonged downlink of data afterwards.

Solar Cells. Solar cells provide a moderately efficient source of power within the inner solar system, converting some 10-20% of the incident solar power into electrical power (depending on cell type, temperature and radiation exposure). As distance from the Sun R increases, the power produced by an array reduces by a factor of approximately R^1.5; the falloff is less then the R² decline in solar irradiance due to the higher efficiency of solar cells at lower temperatures. Solar cells have thus to date been used for missions within the main asteroid belt, although the growing obstacles to the use of RTGs (see below) have led to their use on deep space missions being considered.

For missions in Earth orbit spacecraft often use solar cells in conjunction with secondary batteries to provide power during solar eclipses. Such eclipses are not a problem for deep space probes (except during the parking orbit phase or immediately after trajectory injection), but the use of batteries allows a probe - especially one without sun-tracking arrays - to temporarily repoint itself without losing power. Batteries can also be used to store excess power during quiescent phases of the mission for later use during more intense operations.

Solar cells can either be fixed or sun-tracking. The use of Sun-tracking arrays is more efficient in that it allows the maximum array area to be presented to the Sun at all times. However it involves the use of complex mechanisms and requires that the spacecraft be 3-axis stabilized at all times. If this is not the case then solar cells can be distributed around the body of the spacecraft or on non-tracking deployable panels, but this is a factor of 3 or 4 less efficient in the use of solar cells.

Radioisotope Thermal Generators. RTGs have been used since the 1960s to power satellites for which solar cells are undesirable or impractical. Examples include some military spacecraft for which solar cell arrays would be vulnerable to attack, and outer solar system probes travelling into regions where solar irradiance is very weak. To date, all six probes that have ventured to Jupiter or beyond have been powered by RTGs.

RTGs work by using the heat generated by the radioactive decay of an artificially-produced isotope to heat one end of a solid-state thermoelectric generator [Piscane94], the other end being cooled by exposure to space. The generator consists of series-wired alternate p-type and n-type semiconductor elements; a temperature gradient along each element causes it to produce a potential difference across it due to the thermoelectric (Seebeck) effect. The operating temperature of the hot end is typically 400-600°C and the cold end 100-200°C. Plutonium 238 is commonly used as fuel for RTGs, as its decay half-life of 87 years ensures a fairly constant power flux for even long-duration missions. Shorter missions could use RTGs powered by isotopes with correspondingly shorted half-lives, which (because of the more intense radioactive decay of such isotopes) have higher power/mass ratios. However, such isotopes often have high gamma-radiation fluxes and ould decay during ground storage prior to launch. RTGs are typically small and dense, but in terms of power/mass ratio often compare favourably with solar-cell/battery based systems.

The principle objection to the use of RTGs is the range of environmental and political implications of their use. Flight of RTGs involves extensive design and test work to ensure that any credible launch accident will not lead to the release of radioactive contamination. The shielding required to meet this requirement can add significantly to the mass of the RTG. Furthermore all recent Western RTG launches have met with sustained political objections. RTGs are thus not a suitable candidate power source for this mission.

Power System Design. For this mission the probe will be in sunlight from soon after launch onwards. It was thus decided to use solar cells as the prime power source. However it was recognized that batteries may also be needed at various phases of the mission, depending on the probe configuration:

• Eclipse during parking orbit or following trajectory injection.
• Flyby operations if solar cells insufficient for power demand.
• Data download if extra power required for high-rate data transmission.

To specify the power system in detail requires that the power demand of the spacecraft be known. This was hard to specify in detail during the design process for two reasons: it was very dependent on the probe design and exact power requirements data for different systems is not readily available. Based on the final design baseline though, and using guidelines from Space Mission Analysis and Design and data from previous and proposed small satellite missions [Wertz96, Foquet96] the following estimates were made.

• Communications. For a 5 W transmitted downlink power, SMAD estimates a 30W transmitter power requirement and this is consistent with other missions. It is assumed that the receiver power is much lower than this (sources tended to quote only transmitter supply power) and so 30 W was taken as the communications power requirement.
• Imaging Payload. CCD imagers use relatively little power. Both Clementine and the UoSAT-12 imaging CCDs used about 5-6W each. An estimate of 10W was made to allow a margin for aspects such as filter wheel drive etc.
• On Board Data Handling. Based on SMAD guidelines and the quoted power requirements for the computer and solid-state storage systems on missions such as Ørsted and Freja, 10 W was estimated as the requirement for OBDH.
• Attitude Control System. The two star trackers, based on Clementine data, would use 10 W. Power for the RCS was hard to estimate, as it was highly uneven (e.g. sharp peaks of a few W during thruster actuations) and may involve the need for heaters for thrusters and fuel lines. An average value of 1 W was assumed.

The total power requirement was thus estimated at 61 W. It is possible that this could be reduced for the cruise phase of the mission but it seems likely that all system will be required during and immediately after the target encounter. It should also be noted that this is the power supply required after regulation. Power supply regulators typically dissipate 20 % of the converted power, so the total power required from the solar arrays is 73.2 W.

Solar Array Sizing. The amount of power provided by solar arrays depends on the type of cell used and the distance from the Sun. For this mission it was decided to use gallium arsenide (GaAs) solar cells; although more expensive than traditional solar cells they offer a higher power density. At 1 AU GaAs cells can produce 244 W/m² of unregulated power, which assuming that a peak power tracking converter of 85% efficiency is used will give 207.4 W/m² of regulated power [Larson92]. The probe will be required to operate at up to 1.07 AU from the sun (i.e. 0.05 AU outwards from the Earth at aphelion) at which point the array power will be reduced to 181.2 W/m². To provide 73.2 W thus requires that the probe keep at projected solar array area towards the Sun of at least 0.40 m².

The solar array configuration evolved through several options during the course of this project (see Section 14). The one finally adopted consisted of a cylindrical array spacecraft 1 m in diameter and 0.5 m deep. The cylindrical side face was covered with an 80% array / 20% thermal control surface mix, whilst the front and rear faces each had an annular array with an area half that of the actual face. This meant that each of the front, back and side faces had a square-on projected area to the Sun of 0.40 m². For other Sun angles the total projected area was always greater than this (Fig 15.1)

Battery Sizing. It was decided to provide a primary battery to provide power during any eclipse during the initial parking orbit or early stage of escape trajectory. The worst case parking orbit eclipse is about 40 minutes; eclipse in the early stage of escape trajectory could be considerably longer if the hyperbolic asymptote happened to be along the Earth's shadow cone. Realistically though a trajectory constraint would be imposed to avoid excessively long eclipses for thermal as well as power reasons. It is thus taken that the longest eclipse would be 80 minutes, as this is only slightly longer than the maximum eclipse routinely experienced by GEO satellites (72 minutes). To be conservative and provide the probe's entire regulated power supply for this period would require 97.6 W hr of battery capacity. This could be provided by a lithium thionyl chloride battery of 0.5 kg mass.

Power System Mass. Following the guidelines in SMAD the probe power system mass budget was calculated as in Table 15.1. Note that the solar array masses quoted are for the arrays themselves and do not include the underlying structure, which is discussed in Section 17).