Method and device for deorbiting an artificial satellite from earth orbit by reusing multilayer insulation (MLI)

Abstract

A method and a device for deorbiting artificial satellites with multilayer insulation arranges a deorbiting device beneath it. The device includes a spring element, a bonding agent, and a heating element designed to melt the bonding agent, which at least partially detaches at least one layer of the multilayer insulation from the satellite when a signal is transmitted to the satellite by rolling up or splaying out in the manner of a flap, enlarging its cross-sectional area. The partial detachment exposes the underlying satellite structure to environmental influences. This exposure accelerates the self-disintegration of the artificial satellite, reduces the mass, and increases the ballistic coefficient throughout the deorbiting period. This area enlargement and mass reduction reduce energy, resulting in re-entry into Earth's atmosphere. A number of layers can be arranged so that, regardless of the satellite rotation, at least one surface is always directed against the aerodynamic flow.

Claims

1. A method for deorbiting in a space environment an artificial satellite having a satellite structure and multilayer insulation (MLI), the MLI comprising a plurality of reflective plastic layers coated with metal, each layer being physically disconnected and electrically isolated from adjacent layers by insulating spacer materials, the method comprising: arranging under the MLI a deorbiting device comprising a spring element, a heating element, and a bonding agent, wherein the bonding agent bonds at least one layer of the MLI to the satellite structure, wherein the heating element is designed to melt the bonding agent, and when a signal is actively transmitted to the satellite at least partially peeling off the at least one layer of the MLI from the artificial satellite using the deorbiting device whereby the at least one layer is not separated from the satellite but removed from a part of the satellite by rolling the at least one layer up or by splaying out the at least one layer of the MLI in the manner of a flap to enlarge a cross-sectional area of the artificial satellite via the MLI alone to cause an increase in braking effect of the satellite as well as a reduction in mass due to exposure to the space environment which accelerates deorbiting ultimately.

2. The method according to claim 1, wherein the peeling off of the at least one layer of the MLI takes place at the end of the service life of the artificial satellite.

3. The method according to claim 1, wherein the at least one layer of the MLI is splayed out in the manner of a flap by a spanning force of the spring element.

4. The method according to claim 1, wherein a number of splayed-out layers of the MLI are oriented such that regardless of the position of the satellite, at least a largest surface of one of the splayed-out layers of the MLI is always directed against the aerodynamic flow.

5. A device for deorbiting artificial satellites from Earth orbit according to the method according to claim 1, wherein means are provided for at least partial detachment of the at least one layer of the MLI from the artificial satellite, wherein the means comprises the spring element, the heating element, and the bonding agent, wherein the heating element comprises a thin-film heater, wherein the bonding agent is a heat-soluble adhesive, and wherein the device is configured to be arranged under the MLI of the artificial satellite.

6. The device according to claim 5, wherein the spring element is a flat spring.

7. The device according to claim 5, wherein the spring element is a torsion spring.

8. The method according to claim 1, wherein the at least one layer of the MLI is not separated from the satellite but removed from the part of the satellite by rolling the at least one layer up to expose the part to the space environment.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) Other objects and features of the invention will become apparent from the following detailed description considered in connection with the accompanying drawings. It is to be understood, however, that the drawings are designed as an illustration only and not as a definition of the limits of the invention.

(2) In the drawings,

(3) FIG. 1 shows a satellite according to the invention during the use phase;

(4) FIG. 2 shows a satellite according to the invention after the use phase in a first implementation;

(5) FIG. 3 shows a satellite according to the invention after the use phase in a second implementation;

(6) FIG. 4 shows a device according to the invention for deorbiting in a non-activated state;

(7) FIG. 5 shows a tentering mechanism according to the invention in a non-activated state;

(8) FIG. 6 shows a satellite according to the invention following the use phase with a first activated mechanism; and

(9) FIG. 7 shows a satellite according to the invention following the use phase with a second activated mechanism.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

(10) FIG. 1 shows a satellite (1) according to the invention during the use phase. The device for deorbiting (2) is concealed under the multilayer insulation (3). The multilayer insulation (3) protects the structure and load of the satellite from environmental influences during the use phase, above all from radiation and temperature changes. Following the use phase of the satellite (1), the device for deorbiting (2) according to the invention changes the multilayer insulation (3) to enlarge the surface of the satellite (1). This can take place due to peeling off of the multilayer insulation (3) by means of a peeling mechanism or by spreading the multilayer insulation (3) by means of a tentering mechanism.

(11) FIG. 2 shows a satellite (1) according to the invention following the use phase with activated device for deorbiting, which is formed here as a peeling mechanism (4). On activation of the peeling mechanism (4), the multilayer insulation (3) is rolled up and remains on the satellite (1), which is exposed to radiation exposure in space, whereby the decomposition of the satellite structure, a reduction in mass and ultimately deorbiting commences.

(12) FIG. 3 shows a satellite (1) according to the invention following the use phase with an activated device for deorbiting, which has a tentering mechanism (5). Following peeling off of the multilayer insulation (3), this is extended with the aid of the tentering mechanism (5) in the manner of a flap and thus significantly augments the cross-sectional area of the satellite (1), whereby this is braked sharply.

(13) FIG. 4 shows in detail an inventive peeling mechanism (4) in the non-activated state, consisting of a heating element (6) formed as thin-film heating element, a flat spring (7) integrated in the multilayer insulation (3) and a bonding agent (8).

(14) FIG. 5 shows in detail an inventive tentering mechanism (5) in the non-activated state, consisting of a heating element (6) formed as a thin-film heating element, a bonding agent (8) and a pre-tensioned torsion spring (9) in the multilayer insulation (3).

(15) FIG. 6 shows a satellite (1) according to the invention following the use phase with the mechanism activated after the use phase, consisting of the thin-film heating element (6), the bonding agent (8) and the torsion spring (9), by which the multilayer insulation (3) is extended in the manner of a flap by the spanning force of the torsion spring (9).

(16) FIG. 7 shows a satellite (1) according to the invention following the use phase with the peeling mechanism activated following the use phase, consisting of the flat spring (7), in which the multilayer insulation (3) is rolled up by the tension force of the flat spring (7).

(17) Although only a few embodiments of the present invention have been shown and described, it is to be understood that many changes and modifications may be made thereunto without departing from the spirit and scope of the invention.