GB2125157A

GB2125157A - Spacecraft power system

Info

Publication number: GB2125157A
Application number: GB08318459A
Authority: GB
Inventors: Michael James Baker; Henry William Groves
Original assignee: British Aerospace PLC
Current assignee: BAE Systems PLC
Priority date: 1982-07-08
Filing date: 1983-07-07
Publication date: 1984-02-29
Also published as: GB8318459D0; GB2125157B

Abstract

A power supply system for use on board a spacecraft comprises a Heat engine, a Stirling cycle engine for example, a solar energy concentrator (4, 5) for receiving solar energy and passing it as heat to the hot-side of the engine and a radiator (6) for cooling the cold-side of the engine. The heat engine may produce mechanical energy which is converted by an electrical generator to electrical energy for use on board the spacecraft. <IMAGE>

Description

SPECIFICATION Spacecraft power systems The invention relates to power supplies for spacecraft and is particularly concerned with the provision of an economical method and means for generating electrical power for use by the systems employed in the spacecraft.

Present systems use solar panels employing a multitude of thermo-electric cells which are interconnected and serve to supply electrical power to an on-board battery power supply system.

Not only are such solar panels extremely large which means they must be mounted upon structures which can be held and folded within the compass of the space available within the launch vehicle and therefore are in themselves complicated and expensive, but the materials of the thermo-electric devices degrade with time and therefore the size of the structure and the number of cells has initially to be made over-large for the duty which they are required to meet - thus incurring a severe weight penalty.

The main object of the present invention is to provide an alternative solar powered electrical generating system that is cheaper to produce, relatively lightweight and overcomes these disadvantages.

According to the invention there is provided, on board a spacecraft, a power generating system including a heat engine, a Stirling cycle heat engine for example, with which there is associated a hot-side and a cold-side and which is operable for generating power as a consequence of the transfer of heat between said hot and cold-sides, receiving means for receiving and concentrating solar energy and for ensuring the transfer of the resultant heat energy to said hot-side, radiator means for receiving heat energy from said coldside and emitting it to space, said receiving means and said radiator means being mounted such as to enable the receiving means to be positioned to face the sun while the radiator means faces away from it, and power take-off means coupled to said heat-engine and operable for making the power generated thereby available for use on the spacecraft.

Preferably but not essentially said power takeoff means comprises an electrical power generator for converting mechanical power produced by said engine to electrical power for use on-board the spacecraft.

For a better understanding of the invention, an embodiment thereof will now be described by way of example with reference to the accompanying drawing, in which: figure 1 shows in diagrammatic form a device which operates in a Stirling cycle to drive an electrical generator, and figure 2 shows the arrangement of the heat source and radiator together with device as installed in a satellite as stationed in space.

Referring to figure 1 a thermo-electric generator device 1 which operates on a Stirling cycle utilises diaphragms in conjunction with a regenerator 2 to drive an electrical generator of the form using a soft-iron mass oscillating in a magnetic field.

The operation of the thermo-electric generator is well described in the literature and can be seen in various patents and applications, e.g. British Patent Application 36317/71, Patent Applications Nos. 81-25285 and 81-25284, and British Patents Nos. 2 016 090 and 2 012 886.

In the particular arrangement with which we are concerned, the thermo-electric generator 1 is held to a deployable 2 arm by which it can be stowed, while in the launch phase and brought into the working condition when the spacecraft is placed in orbit.

At the hot end of the regenerator 3 there is attached a solar reflector 4 and a focussing reflector 5 by which solar radiation is collected and focussed. At the cold end of the regenerator 3 there is mounted a heat sink 6 in contact with the "cold finger" or cold space of the regenerator and a diaphragm 7 moving in correspondence with the influence of the heat source and regenerator and that of a return spring 8. A mass of soft iron coupled to diaphragm 7 oscillating within the field of magnet 9 thereby generating electricity which is then fed by line 10 to the spacecraft 1 Oa.

Other arrangements and other devices which utilise such thermal differences which exist in space ranging from 65000C at the hot end to 40K can be used, the normal attitude control means serving to ensure that the reflector is pointed in the required direction.

Within such an arrangement it is envisaged that with the possibility of greater power being made available in smaller space, the hitherto very heavy large capacity batteries may well be discarded and further weight saving be achieved. Furthermore, with greater power available it may be possible to dispense with gas bottle for supplying the thrusters used for attitude control and employ electro-magnetic means to move mass with the spacecraft, thus removing a present restriction on the useful life of a satellite that a limited gas supply presently imposes.

1. A spacecraft comprising a power generating system including a heat engine, a Stirling cycle heat engine for example, with which there is associated a hot-side and a cold-side and which is operable for generating power as a consequence of the transfer of heat between said hot and coldsides, receiving means for receiving and concentrating solar energy and for ensuring the transfer of the resultant heat energy to said hotside, radiator means for receiving heat energy from said cold-side and emitting it to space, said receiving means and said radiator means being mounted such as to enable the receiving means to be positioned to face the sun while the radiator means faces away from it, and power take-off means coupled to said heat-engine and operable

**WARNING** end of DESC field may overlap start of CLMS **.

Claims

**WARNING** start of CLMS field may overlap end of DESC **. SPECIFICATION Spacecraft power systems The invention relates to power supplies for spacecraft and is particularly concerned with the provision of an economical method and means for generating electrical power for use by the systems employed in the spacecraft. Present systems use solar panels employing a multitude of thermo-electric cells which are interconnected and serve to supply electrical power to an on-board battery power supply system. Not only are such solar panels extremely large which means they must be mounted upon structures which can be held and folded within the compass of the space available within the launch vehicle and therefore are in themselves complicated and expensive, but the materials of the thermo-electric devices degrade with time and therefore the size of the structure and the number of cells has initially to be made over-large for the duty which they are required to meet - thus incurring a severe weight penalty. The main object of the present invention is to provide an alternative solar powered electrical generating system that is cheaper to produce, relatively lightweight and overcomes these disadvantages. According to the invention there is provided, on board a spacecraft, a power generating system including a heat engine, a Stirling cycle heat engine for example, with which there is associated a hot-side and a cold-side and which is operable for generating power as a consequence of the transfer of heat between said hot and cold-sides, receiving means for receiving and concentrating solar energy and for ensuring the transfer of the resultant heat energy to said hot-side, radiator means for receiving heat energy from said coldside and emitting it to space, said receiving means and said radiator means being mounted such as to enable the receiving means to be positioned to face the sun while the radiator means faces away from it, and power take-off means coupled to said heat-engine and operable for making the power generated thereby available for use on the spacecraft. Preferably but not essentially said power takeoff means comprises an electrical power generator for converting mechanical power produced by said engine to electrical power for use on-board the spacecraft. For a better understanding of the invention, an embodiment thereof will now be described by way of example with reference to the accompanying drawing, in which: figure 1 shows in diagrammatic form a device which operates in a Stirling cycle to drive an electrical generator, and figure 2 shows the arrangement of the heat source and radiator together with device as installed in a satellite as stationed in space. Referring to figure 1 a thermo-electric generator device 1 which operates on a Stirling cycle utilises diaphragms in conjunction with a regenerator 2 to drive an electrical generator of the form using a soft-iron mass oscillating in a magnetic field. The operation of the thermo-electric generator is well described in the literature and can be seen in various patents and applications, e.g. British Patent Application 36317/71, Patent Applications Nos. 81-25285 and 81-25284, and British Patents Nos. 2 016 090 and 2 012 886. In the particular arrangement with which we are concerned, the thermo-electric generator 1 is held to a deployable 2 arm by which it can be stowed, while in the launch phase and brought into the working condition when the spacecraft is placed in orbit. At the hot end of the regenerator 3 there is attached a solar reflector 4 and a focussing reflector 5 by which solar radiation is collected and focussed. At the cold end of the regenerator 3 there is mounted a heat sink 6 in contact with the "cold finger" or cold space of the regenerator and a diaphragm 7 moving in correspondence with the influence of the heat source and regenerator and that of a return spring 8. A mass of soft iron coupled to diaphragm 7 oscillating within the field of magnet 9 thereby generating electricity which is then fed by line 10 to the spacecraft 1 Oa. Other arrangements and other devices which utilise such thermal differences which exist in space ranging from 65000C at the hot end to 40K can be used, the normal attitude control means serving to ensure that the reflector is pointed in the required direction. Within such an arrangement it is envisaged that with the possibility of greater power being made available in smaller space, the hitherto very heavy large capacity batteries may well be discarded and further weight saving be achieved. Furthermore, with greater power available it may be possible to dispense with gas bottle for supplying the thrusters used for attitude control and employ electro-magnetic means to move mass with the spacecraft, thus removing a present restriction on the useful life of a satellite that a limited gas supply presently imposes. CLAIMS

1. A spacecraft comprising a power generating system including a heat engine, a Stirling cycle heat engine for example, with which there is associated a hot-side and a cold-side and which is operable for generating power as a consequence of the transfer of heat between said hot and coldsides, receiving means for receiving and concentrating solar energy and for ensuring the transfer of the resultant heat energy to said hotside, radiator means for receiving heat energy from said cold-side and emitting it to space, said receiving means and said radiator means being mounted such as to enable the receiving means to be positioned to face the sun while the radiator means faces away from it, and power take-off means coupled to said heat-engine and operable for making the power generated thereby available for use on the spacecraft.

2. A spacecraft according to claim 1 , wherein said heat engine is a Stirling Cycle heat engine.

3. A spacecraft according to claim 1 , wherein said power take-off means comprises an electrical power generator for converting mechanical energy produced by said engine to electrical energy for use on-board the spacecraft.

4. A spacecraft substantially as hereinbefore described with reference to the accompanying drawings.