DE102016224816B4 - Method and device for attitude control of a spacecraft - Google Patents

Abstract

Method for attitude control of a spacecraft (1), in which a force is exerted on the spacecraft (1) by at least one directed beam (7) of electromagnetic radiation of sufficient power for attitude control, which force changes a position and / or orientation of the spacecraft (1) leads in space, wherein the change in position and / or orientation of the spacecraft (1) is brought about by emitting the at least one beam (7) of electromagnetic radiation from the spacecraft (1), characterized in that the electromagnetic radiation with one or a plurality of light emitting diodes (8) arranged on the spacecraft (1) is generated.

Description

Technical field of application

The present invention relates to a method for attitude control of a spacecraft, in which a force is exerted on the spacecraft by at least one directed beam of electromagnetic radiation of sufficient power for attitude control, which leads to a change in the position and / or orientation of the spacecraft in space. The invention also relates to an attitude control device arranged on the spacecraft. Spacecraft can be, for example, artificial satellites, space probes, spaceships, spacecraft, space capsules or space stations.

Various missions of spacecraft planned for use in space, for example satellites or probes, in both the commercial and scientific sectors place increasingly complex requirements on the highly precise stabilizing alignment of the spacecraft or the instruments on board in space. In addition to a fine torque resolution and pulse capability, the fast control-specific responsiveness of the position control system used to correction commands is also decisive. In addition to energy and fuel-optimized generation of control thrusts, a technology that is as contamination-free as possible should often be used in order to avoid influencing other systems of the spacecraft.

State of the art

Known techniques for active attitude control and stabilization use control engines that require fuel to generate a torque of the spacecraft. The force required to generate the torque is conveyed, for example, by oxidation of the fuel educts and subsequent expansion of the combustion gases. An ionized fluid (gas, liquid) can also be accelerated and ejected by means of high electrical fields. In any case, with these techniques, the position-regulating torque generation results from the recoil principle of accelerated, mass-laden fuel.

Another known possibility for position control and stabilization of satellites is based on different variants of reaction or spin wheels. These build up a stabilizing or compensating angular momentum through the adaptive rotational movement of a flywheel. However, they have a maximum value of the angular momentum that they can induce and therefore have to be desaturated from time to time with the aid of thrusters or other angular momentum generators, for example magnetic torques.

Many of the technologies currently in use for active position control of spacecraft use fuel (= mass) to generate torque. This fuel must be carried with you from the start of the mission and is also only available to a limited extent for the duration of the mission. These criteria consequently influence the launch mass and the mission-related service life of the spacecraft. A major disadvantage of these technologies is therefore the need to carry and store the limited fuel available for torque generation.

Even with the initially fuel-free position control by means of reaction wheels, a combination with thrusters as pulse generators cannot be dispensed with in many cases, since the centrifugal masses have to be desaturated when a maximum rotational moment is reached. This also results in greater system complexity. Furthermore, the rotating masses place high demands on the mechanical structure of the spacecraft, since mechanical loads such as oscillations and vibrations have to be absorbed in order, for example, to minimize a negative impact on existing sensors for scientific experiments.

A further disadvantage of these current concepts is that the torque resolution and the reaction times of the position control system, for example due to the starting behavior of reaction wheels, represent limiting factors with regard to the required accuracy. Future planned missions for space-based scientific experiments place requirements for torque in the (sub) micro-newton meter range that cannot be met with existing technology, or only with a great deal of additional effort.

From the DE 198 48 737 C2 a method for position and orbit control of satellites with the aid of electromagnetic radiation is known. In this method, reflection surfaces are formed on the satellite, which are irradiated with a directed beam of electromagnetic radiation from a radiation source located on the ground or on another satellite in order to influence the position and / or orbit of the satellite through the pulse transfer during reflection. Such a technology, however, requires either a line of sight between the earth and the satellite or another spacecraft and, in the case of an earth-based radiation source, is also disturbed by fluctuating losses due to atmospheric influences.

The FR 2 795 457 A1 describes a method for attitude control of a spacecraft according to the preamble of the present claim 1, in which a black body (cavity radiator) on the spacecraft is used to stabilize the spacecraft. The radiation generated by the black body hits a focusing mirror in order to generate directed radiation. For this purpose, the black body is heated using a heating wire.

The GB 2 229 865 A deals with an electric propulsion unit for a spacecraft that works on the basis of a resonant waveguide arrangement.

US 2007/0 045 474 A1 describes a technique for maintaining the position of a swarm of spacecraft relative to one another and therefore requires at least two spacecraft. It is mainly based on the fact that a tether voltage is maintained by counter-blowing with appropriately aligned lasers.

In US 2005/0 001 102 A1, targets on a spacecraft are irradiated with a high-energy beam in order to ablate material there. The ablated material provides a recoil, which is used for position control.

The DE 198 48 737 A1 describes a technique for position and orbit control of satellites, in which high-energy, focused radiation pulses are directed from a ground station onto suitable surfaces of a satellite in order to influence its position and / or orientation.

The object of the present invention is to provide a method for attitude control of a spacecraft and a corresponding device for attitude control on a spacecraft, which enable a highly precise attitude control of the spacecraft and do not require any fuel.

Presentation of the invention

The object is achieved with the method and the device arranged on the spacecraft according to the claims 1 and 8th solved. Advantageous configurations of the method and the device are the subject matter of the dependent claims or can be found in the following description and in the exemplary embodiment.

In the proposed method for attitude control of a spacecraft, a force is exerted on the spacecraft by emitting at least one directed beam of electromagnetic radiation with a radiation power sufficient for attitude control, which leads to a change in the position and / or orientation of the spacecraft in space. The electromagnetic radiation is generated with one or more light-emitting diodes (LEDs) arranged on the spacecraft. The position or orientation of the spacecraft is then changed, if necessary, by emitting the at least one beam of electromagnetic radiation from the spacecraft. The direction of the force acting on the spacecraft is determined by the direction in which the beam is emitted. According to basic physical principles, the magnitude of the force corresponds to 3.3 x 10 ^-9 Newtons per watt of emitted electromagnetic power. To change the orientation of the spacecraft, a torque is generated which essentially results from the recoil impulse of the emitted electromagnetic waves acting via a lever arm. The lever arm corresponds to the distance between the point of emission from a main axis of the spacecraft. With a radiation power of 1 kW with a lever arm of one meter, a torque of 3.3 µNm results.

In the proposed method, the position control is preferably carried out by controlling the direction of emission and / or the emission power of the at least one beam of electromagnetic radiation from the spacecraft. The direction of the torque and thus the change in orientation of the spacecraft can be determined by choosing the direction of emission. The magnitude of the torque exerted in each case can be influenced via the radiated power.

In principle, any type of electromagnetic radiation can be used for the proposed method. A range of electromagnetic radiation is preferably used for which LEDs are available, with which the radiation can be generated with the lowest possible losses and / or as quickly as possible from stored electrical energy.

The radiation can be generated with just one or with several LEDs. The radiation from several LEDs can be bundled into a single beam. Several beams can also be generated, which are emitted, for example, in different spatial directions. The at least one beam of electromagnetic radiation is preferably guided through one or more beam deflection and / or beam guiding elements on the spacecraft before it is emitted. The emission can take place by using one or more movable beam deflection elements in almost any direction or also in fixed predetermined directions via appropriately fixed deflection elements, in which case For example, the beam can be switched between the respective beam deflection elements via beam switches. Of course, other configurations are also possible in order to emit the beam of electromagnetic radiation generated on the spacecraft in different spatial directions. The spatial direction for the radiation is selected so that the desired change in position of the spacecraft is achieved by the radiation.

In a preferred embodiment, the electromagnetic radiation is generated via a plurality of LEDs. The emitted power can then be controlled, for example, by switching one or more of these LEDs on and off. In this way, extremely small thrusts or torques that could not previously be achieved with conventional technology can be generated. This enables the attitude control of spacecraft with extremely high alignment accuracy.

The electrical energy for generating the electromagnetic radiation can be obtained primarily with the aid of photovoltaic cells, for example. In order to compensate for any phases of insufficient solar radiation, rechargeable secondary batteries can also be used. Supercapacitors can also be used for short-term power peaks. Photovoltaics and secondary batteries or supercapacitors are currently preferred but not exclusive options for the primary generation and intermediate storage of electrical power. For example, radioisotope generators, reactors, fuel cells or the like for primary generation and conventional capacitors, regenerative fuel cells or even spin wheels, especially two counter-rotating twist wheels, conceivable for the intermediate storage that may be necessary. Any combination of the above-mentioned components is of course also possible.

The device designed to carry out the method, which is installed on the spacecraft, accordingly has one or more LEDs for generating electromagnetic radiation of sufficient power for position control, an arrangement for forming and / or guiding a beam of electromagnetic radiation and a radiation device , via which the beam of electromagnetic radiation from the spacecraft can be radiated in different directions. The device preferably has a controller for the emission of the beam of electromagnetic radiation for attitude regulation, in particular a controller with which the emission direction and / or the radiation power of the beam of electromagnetic radiation can be controlled in accordance with the desired attitude regulation of the spacecraft.

The proposed method and the associated device use a fuel-free position control, since the torque is generated by emitting electromagnetic waves. Since this type of torque generation does not depend on the conversion of mass-laden fuel, it is possible to dispense with corresponding subsystem components for carrying it along, storing it and converting it to generate torque. This results in a lower mass of the spacecraft, a lower complexity and a longer service life of the attitude control system. This also increases the mission life of the spacecraft. The method and the device enable a shorter reaction time to correction requests in relation to bulk fuel systems whose inherent latency, which is caused by the production, expansion in an engine and possibly by chemical reactions, does not occur in the present method. By using electromagnetic radiation as a pulse generator, very short and extremely precise minimum pulses can be generated that are far superior to the performance specifications of previous position control technologies such as reaction wheels or cold gas engines. The control precision that can be achieved makes very demanding space missions possible. Another advantage is the significantly reduced system complexity due to the elimination of the fuel delivery system. This results in significantly lower costs for the position control system and increased reliability of the position control. Ultimately, this also excludes potential contamination of the spacecraft by backflowing fuel components.

Areas of application of the proposed method and the associated device are all spacecraft which require very precise attitude control. One example is the planned LISA mission of ESA for the detection of gravitational waves, which requires a corresponding fine alignment of laser beams from three satellites over a distance of several million kilometers.

Figure list

• 1 ein Beispiel für einen Raumflugkörper, der eine erfindungsgemäße Vorrichtung zur Lageregelung aufweist; und
• 2 ein Beispiel für die Erzeugung und Abstrahlung des Strahls elektromagnetischer Strahlung von einem Raumflugkörper.

The proposed method and the associated device are explained in more detail below using an exemplary embodiment in conjunction with the drawings. Here show:

• 1 an example of a spacecraft which has a device according to the invention for attitude control; and
• 2 an example of the generation and emission of the beam of electromagnetic radiation from a spacecraft.

Ways of Carrying Out the Invention

1 shows a schematic representation of an example of a spacecraft with a device for attitude control according to the present invention. The spacecraft 1 in this example has photovoltaic modules on its outside 2 as the primary energy source that supplies the electrical components of this spacecraft with energy. For intermediate storage of the photovoltaic modules 2 generated energy is a primary energy store 3 , for example in the form of Li-ion batteries on the spacecraft 1 arranged. In this example, there is an additional energy store for high power requirements 4th provided, which is formed from supercapacitors.

Furthermore is on an outside of the spacecraft 1 a thrust or torque generating unit 5 arranged, which is shown enlarged again in the right part of the figure. This thrust or torque generating unit 5 has a distance R from the main axis 6th of the spacecraft 1 on. The thrust or torque generation unit 5 is part of the device according to the invention and is used to emit a beam 7th electromagnetic radiation from the spacecraft. In the present example, the thrust or torque generation unit enables 5 the radiation of the beam 7th electromagnetic radiation in five spatial directions, which each differ by an angle of 90 ° or 180 °. Depending on the direction of radiation, the spacecraft experiences 1 due to the distance R of the thrust or torque generation unit 5 from the main axis 6th of the spacecraft 1 a torque that changes the attitude of the spacecraft. By controlling the direction of emission and the emission power, it is thus possible to precisely regulate the position of the spacecraft as a function of corresponding control parameters 1 respectively.

2 shows an example of the generation of the beam 7th electromagnetic radiation in a device such as that in the spacecraft 1 the 1 can be used. In this example, an LED array is used as the radiation source 8th used. Such an LED array has a high efficiency. The electrical energy required to generate radiation can be obtained, for example, directly from solar cells or photovoltaic modules, via batteries or, if the power requirement is very high, from supercapacitors (supercaps) to generate high torques, as is also exemplified in 1 are indicated. Via individual glass fibers 9 and a collimator unit 10 In this example, the radiation from the individual LEDs is bundled to form an optical beam 7th . The beam is then in this example via a controllable reflector 11 emitted from the spacecraft. This controllable reflector 11 , which can be controlled mechanically and / or optronically, for example, enables radiation in almost any direction. About the number of activated LEDs in the LED array 8th the generated thrust or torque can be modulated over several orders of magnitude in a simple manner.

List of reference symbols

1

Spacecraft

2

Photovoltaic modules

3

Primary energy storage

4th

additional energy storage

5

Thrust or torque generation unit

6th

Major axis of the spacecraft

7th

beam

8th

LED array

9

Fiberglass

10

Collimator unit

11

Controllable reflector

Claims (11)

Method for attitude control of a spacecraft (1), in which a force is exerted on the spacecraft (1) by at least one directed beam (7) of electromagnetic radiation of sufficient power for attitude control, which force changes a position and / or orientation of the spacecraft (1) leads in space, wherein the change in position and / or orientation of the spacecraft (1) is brought about by emitting the at least one beam (7) of electromagnetic radiation from the spacecraft (1), characterized in that the electromagnetic radiation with one or a plurality of light emitting diodes (8) arranged on the spacecraft (1) is generated. Procedure according to Claim 1 , characterized in that an emission direction of the at least one beam (7) of electromagnetic radiation from the spacecraft (1) and / or an emission power of the at least one beam (7) of electromagnetic radiation from the spacecraft (1) is controlled to carry out the attitude control. Procedure according to Claim 2 , characterized in that the electromagnetic radiation is generated with several light-emitting diodes arranged on the spacecraft (1) and the control of the emitted power is carried out by switching one or more of these light-emitting diodes on and off. Method according to one of the Claims 1 to 3 , characterized in that the beam (7) of electromagnetic radiation is guided through one or more beam deflection and / or beam guiding elements (9, 10) on the spacecraft (1) before it is emitted. Method according to one of the Claims 1 to 4th , characterized in that the electrical energy required to generate the electromagnetic radiation is generated by photovoltaic cells (2) on the spacecraft (1). Procedure according to Claim 5 , characterized in that the electrical energy required to generate the electromagnetic radiation is temporarily stored in rechargeable batteries (3) and / or supercapacitors (4). Method according to one of the Claims 1 to 6th , characterized in that the electromagnetic radiation for the position control is generated with a light emitting diode array. Device for attitude control of a spacecraft (1), which is arranged on the spacecraft (1) and at least - one or more light-emitting diodes (8) for generating electromagnetic radiation of sufficient power for position control, - An arrangement (9, 10) for shaping and / or guiding a beam (7) of the electromagnetic radiation, and - Has an emitting device (5, 11) for emitting the beam (7) of electromagnetic radiation from the spacecraft (1) in different radiation directions. Device according to Claim 8 , characterized in that it has a controller with which the direction of emission of the beam (7) of electromagnetic radiation from the spacecraft (1) and / or an emission power of the beam (7) of electromagnetic radiation from the spacecraft (1) can be controlled for attitude control. Device according to Claim 8 or 9 , characterized in that photovoltaic cells (2) are arranged on the spacecraft (1) with which the electrical energy required for generating the electromagnetic radiation is generated. Device according to Claim 10 , characterized in that rechargeable batteries (3) and / or supercapacitors (4) are arranged on the spacecraft (1), in which the electrical energy required for generating the electromagnetic radiation is temporarily stored.

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