**7.3 Electromagnetic Compatibility (EMC)**

The design of a spacecraft solar array and its power conditioner has to satisfy several requirements, not only in terms of mass, dimensions and power output, but also in terms of electromagnetic compatibility. This is particularly true for scientific mission, when instruments highly sensitive to electromagnetic fields may be boarded. In these cases it becomes crucial for the success of the mission to know which electromagnetic fields are generated at solar array level due to the circulating current and its frequency content, once this is connected to the power conditioning unit. The wires connecting the solar array to the PCDU, via the Solar Array Driving Mechanism (SADM) when necessary, are always twisted pairs (positive and return), but the return connections of the strings are routed on the rear side of the panel, they are not twisted of course, hence the solar array can behave as a transmitting antenna at frequencies which may result incompatible with some of the equipments on board.

Fig. 14. Solar array electrical scheme

These issues are strongly dependent on the power conditioning approach adopted.

In the case of the S3R, with reference to figure 10, it can be seen that within the blue oval there is the shunt switch (MOSFET) together with a linear regulator in order to limit the current spikes at the regulator input when the MOSFET switches ON/OFF. Such spikes are strongly dependent on the total output capacitance of the strings connected in parallel and hence from the capacitance of the single triple junction solar cell. Fewer cells are in a string, or more strings in parallel, higher is this capacitance. The linear regulator can reduce the amplitude of the spikes by a suitable sizing of the dump resistor. For sake of completeness, the inductances present in the circuit diagram are the parasitic ones. Figure 15 shows the frequency spectrum of the current circulating in the harness between solar array and power regulator for different values of the dump resistor. The next figure 16 instead shows the

Architectural Design Criteria for Spacecraft Solar Arrays 177

The direction of the torque is such that the dipole tends to orient itself parallel to lines of

This torque has to be in principle neutralised by the Attitude and Orbit Control System of the satellite, which implies the usage of thrusters (i.e. fuel consumption) or increased authority of magneto-torques and/or reaction wheels (electrical power and mass impact). Clearly there are two ways for the minimisation of this torque; the first one is the minimisation of the areas of the current loops; the second one concern the layout of the solar array strings; adjacent strings can be disposed on the panels in opposite directions, such that the individual torques generated are balanced. With this solution, solar cells having the positive terminal at the string open circuit voltage will lay very close to cells having the

The space plasma is the cause of the accumulation of electrostatic charges on the spacecraft surfaces. The energy of the plasma changes with the altitude; it is around 10,000 eV at about 36,000 km (Geostationary Orbits, GEO) decreasing to 0.1 eV for below 1,000 km (Low Earth Orbits, LEO), within the Van Allen Belts. For what concern the solar arrays it can be said that the interconnections between solar cells and the cell edges are exposed to plasma, and the output voltage resulting at the terminals of a string plays an important role. The worst scenario occurs at BOL, at the minimum operative temperature (eclipse exit). In these conditions the open circuit voltage is at the maximum value, if triple junction solar cells are used and a string is for instance composed of 34 cells, this voltage can be above 90V; this is

The value of the maximum current that can flow through a conductive part of the array (usually the current of a single string if each is protected by a diode) is also important; indeed it has been proofed that in order to have a self sustained secondary arc, minimum value of the current for a particular voltage is needed. In case of ECSS standard applies, in particular "Spacecraft Charging – Environment Induced Effects on the Electrostatic Behaviour of Space Systems (**ECSS-E-20-06**)", then it can be said that no tests are required to prove the safety of the solar array to secondary arcing when the maximum voltage-current couple available between two adjacent cells on the panel, separated with 0.9mm as nominal

An inter-cell gap between strings of adjacent sections may be defined at 2 mm, cell to cell, that means 1.85 mm between cover-glasses. Finally, taking into account tolerances of the tools used during manufacturing of the solar array, it results that the distance between

70 V 0.6 A No self sustained secondary arcing possible 50 V 1.5 A No self sustained secondary arcing possible 30 V 2 A No self sustained secondary arcing possible 10 V - Voltage is too low to allow any arcing

force of *B*, minimizing the potential energy and achieving a stable position.

negative terminal at 0V. And this opens the door to another issue to be faced.

(17)

sin *T M B MB*

**7.5 Electrostatic Discharges (ESD)** 

the maximum voltage between two adjacent cells.

value, is below the threshold in the following table:

**VOLTAGE CURRENT COMMENTS** 

adjacent strings is always higher than 1.6 mm

Table 2. ESD limit conditions

frequency spectrum of the current for the same solar array section when the power conditioning is made by a buck converter with a MPPT control loop. It can be immediately seen that in case of MPPT power conditioning the current ripple on the solar array harness is much lower at low frequencies, not higher than 8 mA; and therefore such solution may be interesting when the power subsystem has to cope to very stringent requirements from EMC point of view.

Fig. 15. Frequency spectrum of Solar Array output current for S3R power conditioning

Fig. 16. Frequency Spectrum of Solar Array output current for MPPT power conditioning

#### **7.4 Effect of the Earth magnetic field**

The interaction between the Earth magnetic field *B* and the currents circulating in each string generate a torque disturbing the desired attitude of the whole spacecraft. The magnetic moment *M* due to the current is given by

$$M = I \cdot A \tag{16}$$

Where *I* is the current and *A* is the area of the current loop; in the case of the solar array this area corresponds in a first approximation to cross section of the panel substrate; on the front face of it the cells are mounted, on the rear face the return harness is implemented. The resulting torque is

frequency spectrum of the current for the same solar array section when the power conditioning is made by a buck converter with a MPPT control loop. It can be immediately seen that in case of MPPT power conditioning the current ripple on the solar array harness is much lower at low frequencies, not higher than 8 mA; and therefore such solution may be interesting when the power subsystem has to cope to very stringent requirements from

Fig. 15. Frequency spectrum of Solar Array output current for S3R power conditioning

 Frequency 183Hz 300Hz 1.0KHz 3.0KHz 10KHz 30KHz 80KHz

Fig. 16. Frequency Spectrum of Solar Array output current for MPPT power conditioning

Frequency 0Hz 5MHz 10MHz 15MHz I(R\_SA)

The interaction between the Earth magnetic field *B* and the currents circulating in each string generate a torque disturbing the desired attitude of the whole spacecraft. The

Where *I* is the current and *A* is the area of the current loop; in the case of the solar array this area corresponds in a first approximation to cross section of the panel substrate; on the front

face of it the cells are mounted, on the rear face the return harness is implemented.

*M I A* (16)

**7.4 Effect of the Earth magnetic field** 

100fA 1.0pA 10pA 100pA 1.0nA 10nA 100nA 1.0uA 10uA 100uA 1.0mA 10mA 100mA 1.0A 10A 100A

The resulting torque is

magnetic moment *M* due to the current is given by

I(L\_harness)

0A

40mA

80mA

120mA

EMC point of view.

$$T = M \times B = M \cdot B \cdot \sin \mathcal{G} \tag{17}$$

The direction of the torque is such that the dipole tends to orient itself parallel to lines of force of *B*, minimizing the potential energy and achieving a stable position.

This torque has to be in principle neutralised by the Attitude and Orbit Control System of the satellite, which implies the usage of thrusters (i.e. fuel consumption) or increased authority of magneto-torques and/or reaction wheels (electrical power and mass impact).

Clearly there are two ways for the minimisation of this torque; the first one is the minimisation of the areas of the current loops; the second one concern the layout of the solar array strings; adjacent strings can be disposed on the panels in opposite directions, such that the individual torques generated are balanced. With this solution, solar cells having the positive terminal at the string open circuit voltage will lay very close to cells having the negative terminal at 0V. And this opens the door to another issue to be faced.

#### **7.5 Electrostatic Discharges (ESD)**

The space plasma is the cause of the accumulation of electrostatic charges on the spacecraft surfaces. The energy of the plasma changes with the altitude; it is around 10,000 eV at about 36,000 km (Geostationary Orbits, GEO) decreasing to 0.1 eV for below 1,000 km (Low Earth Orbits, LEO), within the Van Allen Belts. For what concern the solar arrays it can be said that the interconnections between solar cells and the cell edges are exposed to plasma, and the output voltage resulting at the terminals of a string plays an important role. The worst scenario occurs at BOL, at the minimum operative temperature (eclipse exit). In these conditions the open circuit voltage is at the maximum value, if triple junction solar cells are used and a string is for instance composed of 34 cells, this voltage can be above 90V; this is the maximum voltage between two adjacent cells.

The value of the maximum current that can flow through a conductive part of the array (usually the current of a single string if each is protected by a diode) is also important; indeed it has been proofed that in order to have a self sustained secondary arc, minimum value of the current for a particular voltage is needed. In case of ECSS standard applies, in particular "Spacecraft Charging – Environment Induced Effects on the Electrostatic Behaviour of Space Systems (**ECSS-E-20-06**)", then it can be said that no tests are required to prove the safety of the solar array to secondary arcing when the maximum voltage-current couple available between two adjacent cells on the panel, separated with 0.9mm as nominal value, is below the threshold in the following table:


Table 2. ESD limit conditions

An inter-cell gap between strings of adjacent sections may be defined at 2 mm, cell to cell, that means 1.85 mm between cover-glasses. Finally, taking into account tolerances of the tools used during manufacturing of the solar array, it results that the distance between adjacent strings is always higher than 1.6 mm

Architectural Design Criteria for Spacecraft Solar Arrays 179

pointing attitude of the instruments results in highly variable illumination of the panel, therefore the computation energy budget can be quite challenging because the power subsystem may have power coming from both solar array and battery pack at the same time along the orbit. This behaviour may significantly reduce the useful time for the recharge of the battery in sunlight, and an oversized solar panel may be needed. The ESA spacecraft GOCE is a good example of such body mounted panels; two of them are installed on the fixed "wings" of the satellite, the other two are on the "fuselage". It is worth to note that the temperatures on the solar panels are very different between one another, this because of the different illumination levels and different thermal exchange of the wings (remaining colder) with respect to the fuselage (hotter panels). Such configuration, dictated by many other requirements at satellite level, can have a huge impact in the complexity of the power

The third one is the classical double deployable wing. This solution is classical for telecommunication geostationary satellites. Each wing is moved by a Solar Array Driving Mechanism having the rotation axis perpendicular to the orbital plane. The illumination is optimized by the automatic orientation of the panels. This kind of configuration is the best solution when several kilowatts are needed, as in the case of recent telecom satellites. Each wing is then composed of several panels kept folded at launch, and then progressively deployed by suitable mechanisms at early phase of the mission. The satellite Hylas-1 gives a

The following two examples will show how a spacecraft solar array, composed of one or more panels having different orientations, provide the needed power during the mission. The examples reported consider body mounted panels, having a fixed orientation with respect to the satellite body axes. This kind of panels is typically used for small and medium satellites, with a power demand less than 1 kW. If on one hand they are relatively cheap and easy to realise, on the other they may require additional efforts for the proper assessment of the energy budget throughout the orbits. This is particularly true in case of the power bus is an unregulated one, having wide voltage variations because of the battery

conditioning concept to be adopted.

good example for such solution.

Fig. 19. Deployable Solar panels, Hylas-1 (Credits: ESA - J. Huart.)

**9. Design and simulation examples** 

**8.3 Deployable wings** 
