**6. The power and energy budget**

The starting point for the solar array sizing is the correct identification of the power demand throughout the whole mission of the spacecraft.

Such power demand may change during the satellite lifetime either because of different operational modes foreseen during the mission or, more simply, because of degradation of the electrical performances of the electrical loads (in majority electronic units).

Taking into consideration what just said, an analysis of power demand is performed, including peak power, of all the loads installed either in the platform or as payload for each identified phase of the mission. Because of presence of sun eclipses, and possible depointings along the orbit, an analysis of the energy demand is also performed, this because in case of insufficient illumination the on board battery will supply the electrical power, and the solar array has to be sized in order to provide also the necessary power for its recharge. The power budget is based on peak power demands of the loads, while the energy budget is based on average consumptions.

It is good practice consider power margins both at unit and electrical system level.

The consumption of each unit is calculated considering the following criteria:


Several electronic units work in cold or hot redundancy; this has to be taken into account when summing the power demands.

Once the power demand is defined including the margins above, it is advisable to add 20% extra margin at system level and defined at the beginning of the project. Such margin is particularly useful during the satellite development in order to manage eventual power excesses of some units beyond the margins defined at unit level. In this way eventual Request For Deviation (RFD) issued by the subcontractors can be successfully processed

Architectural Design Criteria for Spacecraft Solar Arrays 173

conditioning concept, and some sizing constraints mainly raised by the space environment

The first concept is based on the use of a shunt regulator; the figure below shows the electric schematic of a cell of a Sequential Switching Shunt Regulator (S3R), several solar array strings can be connected in parallel to the input of the regulator's cell; the voltage at the terminals of the output capacitor (Main Bus capacitor) is regulated by the switching of the

**R)** 

such as electrostatic discharges and earth magnetic field.

MOSFET contained in the blue oval.

**7.1 Regulation based on Sequential Switching Shunt Regulator (S3**

Fig. 10. Electrical Section of a Sequential Switching Shunt Regulator (S3R)

Fig. 11. Solar array working points as function of required power

5

10

Current [A]

15

The operating voltage of the solar array is constant and equal to main bus nominal voltage plus the voltage drops due the two diodes in series along the line, the solar array harness, and the blocking diode placed at the string positive output. In case of a fully regulated power bus, this operating voltage remains fixed during both sunlight and eclipse periods throughout the orbit; if the power bus is instead a battery regulated one it implies that the bus voltage decreases during eclipse periods, when the battery discharges, provoking a

<sup>0</sup> <sup>10</sup> <sup>20</sup> <sup>30</sup> <sup>40</sup> <sup>50</sup> <sup>60</sup> <sup>70</sup> <sup>80</sup> <sup>90</sup> <sup>100</sup> <sup>0</sup>

Voltage [V]

**Solar Array Performances with S3R**

**Demanded current at eclipse exit** 

solar array Hot 18s-20p solar array Cold 18s-20p solar array Hot 18s-25p solar array Cold 18s-25p power curve 280W power curve 320W

**Available power for 18s-25p at eclipse exit** 

Supposing a power need of 280W, Figure 11 shows that a solar array composed of 20 strings of 18 cells (18s – 20p), at the eclipse exit (Varray= 27V) cannot provide the required power. In this condition the battery keeps discharging, lowering further down the operating voltage. This power bus lock-up has to be avoided increasing the number of strings in parallel. Adding 5 more strings (i.e. 25% more) the solar array can deliver 320W at 27V when cold;

migration of the operating point of the solar array towards the short circuit one.

without endangering the whole spacecraft design. This is particularly true for scientific missions, where many times the development of the instruments may reveal so challenging that an excess of power demand cannot be excluded a priori.

At this point harness distribution losses are introduced, 2% of the power demand defined with all margins at unit and system level may be a good compromise between losses containment and harness mass.

The Power Control and Distribution Unit (PCDU) is the electronic unit devoted for the solar array and battery power conditioning and regulation, power distribution and protection, execution of received telecommands (i.e. switch on/off of the loads) and telemetry generation. Its power consumption without considering the efficiencies of primary bus power converters depends on the management of the digital interfaces with the on-board computers, the control loop and protection electronics, the value of such consumption is not immediate to calculate but it can be said that a PCDU capable to manage 1kW can consume about 30W. However it consumption strongly depend on the number of implemented distribution lines, and relevant electronic protections.

Now its time to add the power needed for the recharge of the battery, this power strongly depend on the mission profile, and many times the maximum discharge of the battery occurs at launch, from lift-off up to the successful sun acquisition by the satellite with optimal sun pointing of the solar panels. Some times due to the complexity of the satellite design and mission profile it is not possible to have a full recharge of the battery in one orbit before the next eclipse, then the power allocated for such incumbency has to assure a positive battery recharge trend throughout a limited number of orbits.

The power delivered by the solar array is conditioned by suitable power converters in order to provide it to the loads with a regulated voltage, or at list with the voltage varying between a maximum and minimum value. These converters may have an efficiency between 98.5% and 95% and the choice of their topology is made according to several criteria and constraint dictated by the overall satellite system design. Such efficiencies are taken into account adding up to an additional 5% to the budget defined so far.

The harness losses between solar array and PCDU may be calculated having as objective 1V voltage drop at the maximum required power; again, considerations about the harness mass can provoke the change of such objective.

Finally, in case of the European ECSS standard (ECSS-E-ST-20C) is considered as applicable, an additional 5% margin on power availability shall be assured at the satellite acceptance review End of Life (EOL) conditions and one solar array string failed.
