GB2067034A - Power distribution arrangement - Google Patents

Abstract

In a spacecraft, outputs from respective solar power panel sections (S1-SN) are connected via respective isolating diodes (6) to respective load buses (B1-BN) and also via cross-connect diodes (7) to others of the load buses so that each bus can receive power from at least two panel sections. A switchable standby power supply (8-11) such as one or more storage batteries may be connected to the load buses downstream of the diodes. <IMAGE>

Description

SPECIFICATION A spacecraft power supply and distribution system The invention relates to improvements in power supply systems for spacecraft, e.g. satellites and in particular to power supplies incorporating solar cells possibly together with a standby source such as a battery.

The power requirements of present satellites often means that a single battery consisting of single cells in series is no longer capable of supplying sufficient energy throughout an eclipse period to support the satellite payload.

To support the payload more than one battery is required. If all the payload is supplied from one busbar, batteries therefore have to be connected in parallel directly to the busbars, or via convertors. In the former alternative, attendant problems of power sharing between the batteries result.

One solution is to split the payload between more than one busbar. Rarely is it possible to make this split equal, and thus one bus will inevitably supply more power than the other.

For a battery, which is capable of supply of high powers but is energy limited, the result will be a higher depth of discharge than can usually be allowed, since batteries are normally only discharged to about 60% of their actual capacity.

However, for solar array which is power limited, array area must be sized to meet the maximum power demand of the payload.

Sometimes the payload requirements alter throughout life, or there are occasional pulse loads. To accommodate these loads by splitting the solar array completely to form two or more power systems usually results in arrays whose cumulative area is greater than the unsplit array.

According to this invention, there is provided, in a spacecraft, an electrical power supply and distribution system comprising: solar power generating means having a plurality of outputs, a plurality of load buses for supplying power to respective load groups, for each said generating means output, a first and at least one more unidirectional energy flow device connected between the associated output and respective ones of at least two of said load buses, each load bus being thereby connected via one of the first unidirectional energy flow devices to one of said outputs and via a respective unidirectional energy flow device to at least one other of said outputs, and the different load buses being connected to respective different combinations of at least two of the outputs.

Advantageously, switchable standby power supply means are connected to said load buses downstream of said unidirectional energy flow devices.

For a better understanding of the invention and to show how the same may be carried into effect, reference will now be made, by way of example, to the accompanying drawings, in which: Figure 1 is a simplified diagram illustrating the principles of the invention, and Figure 2 and 3 are respective satellite power supply and distribution systems incorporating the principles of Fig. 1.

The following description is concerned with a satellite power system which has a multiple busbar structure and which utilises solar arrays and batteries as power sources. While distribution of power between busbars is a simple matter during eclipse, and can be made near equal, in sunlight the distribution changes and becomes varied and unequal.

When the size of the solar array is strictly limited it is not easy to include area contingency on the split array sections for load diversity. Battery connection to the payload may have to be for eclipse periods only and connection may have to made without discontinuity of supply.

As shown in Fig. 1, a spacecraft power supply and distribution system may comprise a plurality N of solar power generating panels or panel sections S1, S2, S3-SN each having an output line 5, and a corresponding plurality of load buses B1, B2, B3-BN each coupled to a respective load group. Each output line 5 is connected via a first recitifier diode 6 to a respective one of the load buses and via a second rectifier diode 7 to another of the load buses such that the different load buses are connected via respective first and second diodes 6 and 7 to respective different combinations of two of the lines 5. Thus the output line of panel S1 is connected via a first diode 6 to bus B1 and via a second diode 7 to bus B2, the output line of panel S3 is connected via a first diode 6 to bus B2 and via a second diode 7 to bus B3 and so on.

The system further includes two standby batteries 8 and 9. Battery 8 is connected via respective on-off switches 10 to each of some of the buses downstream of the recitifier diodes 6 and 7 while battery 9 is connected via respective on-off switches 11 to each of the remaining buses again downstream of the rectifier diodes. In the figure only the buses B1, B2, B3 and BN are actually shown and so battery 8 is shown connected via two switches 10 to buses B1 and B2 while battery 9 is shown connected via two switches 11 to the remaining ones of the shown buses, i.e. B3 and BN. It will be realised however, that battery 8 would probably be connected via an appropriate number of switches to say half the number N of buses while battery 9 would be connected to the other half.It will be realised also that the particular arrangement is modifiable-for example there could be only a single standby battery and switches connecting it to all the buses or there could be more than two standby batteries.

If it is assumed that the voltages Vb,,s on all the load buses are equal, then the total power output P supplied by all the solar panels to all the load is given by the expression:

where ILn is the load current drawn by the nth bus.

Also, P = Vbu l, x N where I,, is the current supplied by each panel. The latter equality is true because of the second or "cross" diodes 7 which allow unused solar panel power on one bus to be directed to another, provided that

i.e. provided that the power budgetary demands on successive load buses do not greatly exceed the power output of a single solar panel or section and that they tend to be successively distributed about the single panel output capability.

The diodes 7 tends to equalise the bus voltages within the variation of voltage drops across all the diodes.

When no power is delivered to the buses from the solar arrays the batteries are switched to the load buses to supply power and the diodes then block reverse flow of current into the solar array panels.

By splitting the solar array sections in the first instance, the power system is made less vulnerable to total failure in the event of a short circuit on one bus bar. In the system shown in Fig. 1, a short circuited busbar would reduce the power output of the system by the output power of two array sections only and all other busbars would be available for use, albeit within the power capability of the remaining (n-2) sections.

As shown in Fig. 2, in a more practicable realisation of the principle demonstrated with reference to Fig. 1, the number of separate solar panels or panel sections is double the number of load buses and each two panels or sections are connected via respective pairs of first and second diodes to the same two load buses. In Fig. 2, there are six load buses B1 to B6 and correspondingly six solar panels or panel sections S1 to S6 but, in addition, six more panels S11 to S16-the panels being shown symbolically as current generators.

Each of the panels S1 to S6 is connected, correspondingly as in Fig. 1, by way of a first rectifier diode 6 to one of the load buses and by way of a second rectifier diode 7 to another load bus such that the respective load buses receive feeds from respective different combinations of two of the panels S1 to S6.

Each of the panels S1 1 to S16 is also connected by way of a first rectifier diode 6 to one of the load buses and by way of a second diode 7 to another bus. Thus, the panels S1 and S11 are connected via respective diodes 6 to the bus B1 and via respective diodes 7 to the bus B6, the panels S2 and S12 are connected via respective diodes 6 to the bus B1 and via respective diodes 7 to the bus B2 and so on. This arrangement gives greater reliability because, if there is a failure of say the panel S1 or the diode 6 through which panel S1 is connected to load bus B1, then the load bus B1 can still receive some power from panel S11 and from panels S2 and S12.

The Fig. 2 system is again provided with two standby batteries 8 and 9 but here each battery is connectible to any one or more of the load buses by way of selector switches SW in such a way that any of the buses may be fed from either battery or some may be fed from one battery while other buses are fed from the other battery. In addition each battery has a centre tap i.e. a connection made to between two intermediate cells of the battery, and these taps are connected via respective rectifier diodes 21 and 22 to a common line 23 which is in turn connected via respective diodes 24 to 29 to respective ones of the load buses downstream of the switches SW.

In the power system of Fig. 3, additional reliability engendering component redundancy and other measures are incorporated. Here there are four main load buses B1 to B4 and each of these is fed via three first diodes 6 from three of the solar panel sections S1 to S12 and via three second diodes 7 from three more of the panel sections S1 to S12. Thus, as in Figs. 1 and 2, each of the panel sections S1 to S12 is connected via a first diode 6 to one of the panels B1 to B4 and via a second diode 7 to another of these panel sections.

That side of each of the sections S1 to S12 which is connected to diodes 6 and 7 is also connected to one side of a respective one of twelve further panel sections S21 to S32. The other side of each of the sections S2, S5, S8, S10 and S11 is connected through one of five more panel sections S42, S45, S48, S50 and S51 to a common ground line 60 while the other sides of the panel sections S1, S3, S4, S6, S7, S9 and S12 are connected directly to this line. Connections are made from the points of interconnection between panel sections S2 and S42, S5 and S45, S8 and S48 and S11 and S51, and from the diode side of sections S1 and S7 through a selector switch arrangement 61 to the line 60.The selector switch arrangement 61, which itself comprises switching path redundancy, enables selection and/or isolation of various panel sections.

Two standby batteries 8 and 9 are provided, as in Fig. 2, connected via selector switches SW such that any bus can be fed from either battery. Also, as in Fig. 2, a centre-tap from each battery is connected via a diode 21 or 22 to a common line 23 which is in turn connected to each bus via a respective diode 24 to 27. Here, however, each diode 21, 22 and 24 to 27 is connected in parallel with another diode to give redundancy.

Those sides of the panel sections S21 to S3 which are remote from the panel sections S1 to S12 are connected via rectifier diodes 62 and swiiches 63 to the selector switches SW such that the buses are able to receive power via a route other than those containing diodes 6 and 7.

Each of the load buses is connected via switch arrangements 64, which include redundant switch paths, and sub-buses 65 to various load groups, the switch arrangements enabling feeds to the different load groups from different ones of the bus lines B1 to B4.

It will be appreciated that the system of Fig.

3 has been described mainly with a view-to illustrating the application of the principles disclosed herein to a practicable spacecraft power system and that the particular circuit design shown, particularly the degree of component and supply path redundancy provided, is adapted in practice to a particular spacecraft according to factors such as the priority or importance of particular loads and groups thereof.

It will be further appreciated that although the terms solar panel and/or panel section have been used herein, this does not necessarily imply any clearly separate elements in a physical sense. Generally a spacecraft will comprise one or more solar arrays, for example, it might comprise two wing-like structures extending out from the craft when deployed and each having a large number of solar cells mounted thereon. Simply as a matter of mechanical design, it may well appear that these cells are grouped into panels or sections in the physical sense. However, the term panel section used herein refers simply to a set of interconnected ones of the cells however physically arranged.

Claims (6)

1. In a spacecraft, an electrical power supply and distribution system comprising: solar power generating means having a plurality of outputs, a plurality of load buses for supplying power to respective load groups, for each said generating means output, a first and at least one more unidirectional energy flow device connected between the associated output and respective ones of at least two of said load buses, each load bus being thereby connected via one of the first unidirectional energy flow devices to one of said outputs and via a respective unidirectional energy flow device to at least one other of said outputs, and the different load buses being connected to respective different combinations of at least two of the outputs.

2. A spacecraft power supply and distribution system according to claim 1, including switchable standby power supply means connected to said load buses downstream of said unidirectional energy flow devices.

3. A spacecraft power supply and distribution system according to claim 2, wherein said switchable standby power supply means comprises at least one electrical storage battery and switch means connected between the battery and said load buses.

4. A spacecraft power supply and distribution system according to claim 1, 2 or 3 including a further plurality of outputs from said solar power generating means and further unidirectional energy flow devices connected between each further output and said load buses.

5. A spacecraft power supply and distribution system according to claim 1, 2, 3 or 4 wherein said unidirectional energy flow devices comprise rectifier diodes.

6. In a spacecraft, an electrical power supply and distribution system substantially as hereinbefore described with reference to Fig.

1, 2 or 3 of the accompanying drawings.