JP2013233906A - Spacecraft - Google Patents

Abstract

PROBLEM TO BE SOLVED: To solve such a problem that, when the conventional power generation method using a solar panel or the like is applied in a spacecraft that searches an extraterrestrial planet or the like, the efficiency of power generation per unit area of a solar cell is lowered because of low solar light intensity, and therefore, the mounting of a large-size solar cell panel is required and questions are caused such as an increase in mass of a spacecraft and the shortening of design life.SOLUTION: Surface treatment etc. for collecting solar light is performed for an opening surface of a communication antenna attached to a spacecraft from the beginning, and the solar light with which the communication antenna is irradiated is collected to a solar cell attached to a lateral surface of the spacecraft. An amount of power generation per unit area of the solar cell is thereby improved and the mounting of a solar cell panel is made unnecessary and as a result, the effect of reducing mass of the spacecraft and of improving temperature control within the spacecraft is obtained.

Description

  The present invention relates to a spacecraft such as an artificial satellite equipped with a communication antenna having a function of collecting sunlight. More specifically, the spacecraft panel is attached to the outer surface of the spacecraft panel to reduce the size and weight of the spacecraft, and efficiently exhaust the excessive heat input of the concentrated sunlight outside the spacecraft. Thus, the present invention relates to a spacecraft that improves the efficiency of temperature control.

  In a spacecraft operated on a geostationary orbit or low altitude orbit, power generation is generally performed using a solar cell. The spacecraft attitude control method mainly includes a spin stabilization method and a triaxial stabilization method, and the solar cell mounting method differs depending on these methods.

  In the spin stabilization method, a solar cell is directly attached to the outer peripheral surface of a cylindrical space mechanism. Since sunlight is incident from the side of the cylindrical space mechanism body, only about half of the attached solar cells are simultaneously irradiated. In addition, the amount of power generated is limited by the size of the space mechanism to which the solar cell is attached, and there is a problem that sufficient power cannot be secured.

  On the other hand, in the triaxial stability method, a solar cell panel in which a large number of solar cells are mounted on a plane facing the irradiation direction of sunlight from a rectangular parallelepiped space mechanism panel. The solar cell panel is provided with a drive mechanism so as to always face the sun. Thereby, it becomes possible to always receive sunlight on the entire surface of the solar cell panel, and sufficient electric power can be secured. This method is the mainstream of current spacecraft power generation systems.

  The amount of power generated by the solar cell is proportional to the installation area and the intensity of sunlight. Since the sunlight intensity is determined by the positional relationship between the sun and the spacecraft, the solar cell panel area is adjusted according to the amount of power generation required for the spacecraft.

  In spacecraft that orbit the earth, the solar panel is small because the spacecraft and the sun are close to each other and the sunlight intensity is high.

  However, when exploring extraterrestrial planets and the like, the distance from the sun increases, so the spacecraft is placed in an environment with low sunlight intensity. That is, the amount of power generation per unit area is lower than that of a spacecraft orbiting the earth. Taking a Jupiter spacecraft as a specific example, the sunlight intensity near Jupiter is about 4% near the Earth, and the power generation amount per unit area of the Jupiter spacecraft is about 4% of the spacecraft orbiting the Earth. As a result, in order to supply the same amount of power as the spacecraft orbiting the earth, it is necessary to increase the size of the solar cell panel by about 25 times.

  In order to solve the problem of an enlarging solar cell panel, Patent Document 1 discloses a technique for attaching a reflecting mirror around the solar cell panel and condensing sunlight on the solar cell panel. The technology of Patent Document 1 has a configuration in which all the sunlight irradiated to the reflecting mirror is collected on the solar cell panel, and the area of the solar cell panel can be reduced to about half.

  Patent Document 2 discloses a technique for generating power by concentrating sunlight irradiated onto a hemispherical reflecting mirror on a solar cell disposed near the focal point of the reflecting mirror. Although this technology is not intended to be applied to spacecraft, the power generation efficiency of solar cells can be increased by using this technology, and the number of solar cells can be greatly reduced. .

Japanese Patent Laid-Open No. 11-245897 JP 2011-129847 A

When exploring extraterrestrial planets and the like in this way, the distance from the sun is far and the intensity of sunlight is weak. When the conventional method is applied in an environment where the sunlight intensity is weak, it is necessary to attach a huge solar cell panel in order to obtain electric power necessary for the operation of the spacecraft. When a huge solar cell panel was installed, there was a problem that the mass of the spacecraft increased.
The increase in mass increases the amount of propellant used when performing attitude control and orbit control of the spacecraft, and there is a problem that the design life of the spacecraft is shortened.

Conventionally, as described in Patent Document 1 and Patent Document 2, a technique for combining a solar cell panel with a reflecting mirror or the like has been disclosed for such a problem. In these techniques, when a reflecting mirror or the like is not used, it is necessary to attach a reflecting mirror having an area equivalent to or larger than the required area of the solar cell. That is, in the techniques of Patent Document 1 and Patent Document 2, the area of the solar cell can be reduced, but a large-sized deployable structure is still necessary for power generation, and the above-described problems cannot be solved.
Further, a part of the sunlight absorbed by the solar cell is converted into electric power, but the remaining sunlight is converted into heat. When sunlight is collected as in Patent Document 1 and Patent Document 2, the amount of power generation per unit area increases, but the amount of heat input by sunlight also increases, and the temperature of the portion where sunlight is collected rises. To do. Due to the high temperature of the solar cell, other problems such as deterioration of the performance of the solar cell and failure of the electric circuit occur. As described above, when sunlight is collected in a spacecraft to generate electric power, excessive heat input occurs as another problem.

  The present invention has been made to solve such problems, and has an object to provide a solar concentrating communication antenna capable of reducing the weight of a spacecraft by reducing the number of solar battery panels mounted on the spacecraft. To do.

  A spacecraft according to the present invention is a spacecraft provided with a communication antenna, and a structure panel constituting the spacecraft includes a solar cell that receives sunlight collected by the communication antenna.

  According to the sunlight collecting communication antenna according to the present invention, sunlight can be collected while being used as a communication antenna as in the conventional case. Thereby, a solar cell panel can be reduced and the mass of a spacecraft can be reduced without adding a new structure.

It is the schematic which shows the structure of the sunlight condensing communication antenna which concerns on Embodiment 1 of this invention. It is a fragmentary sectional view of the sunlight condensing communication antenna which concerns on Embodiment 1 of this invention.

Embodiment 1 FIG.
FIG. 1 is a schematic diagram of a configuration of a spacecraft such as an artificial satellite according to Embodiment 1 of the present invention. A communication antenna 3 that has been subjected to a surface treatment or the like that increases the solar reflectance is attached to the space mechanism panel 1a (hereinafter simply referred to as the structure panel 1a).
For example, the communication antenna 3 has a function of reflecting and condensing sunlight in addition to a communication function of communicating with a ground station. The communication antenna 3 is deployed from the structure panel 1a and is arranged so as to face the earth.
The solar cell 2 is attached to the structure panel 1a at a position that becomes the focal point of the communication antenna 3. The sunlight 4 irradiated to the opening surface of the communication antenna 3 is reflected and condensed on the solar cell 2 attached to the structure panel 1a to generate power.

  For the communication antenna 3, a parabolic antenna, a spherical antenna, or the like is used so that both radio waves and sunlight 4 can be reflected. The antenna opening surface is subjected to a surface treatment such as mirror finishing in order to increase the solar reflectance. For example, a coating process is performed on the surface of the antenna opening in order to increase the sunlight reflectance. Alternatively, a material that reflects radio waves and has a high sunlight reflectance may be attached to the antenna opening surface.

  FIG. 2 is a partial sectional view showing the configuration of the embodiment according to the present invention. In order to diffuse the heat generated in the solar cell 2 into the spacecraft, a heat pipe 6a having a high heat transport capability is embedded in the structure panel 1a. In addition, the variable emissivity element 5 is attached to the space mechanism panel 1b (hereinafter referred to as the structure panel 1b) in order to dissipate heat generated by the internal device 7 and heat generated by the solar cell 2. The heat pipe 6b is also embedded in the structure panel 1b.

  The heat pipe 6a and the heat pipe 6b are connected to enhance the thermal coupling between the structure panel 1a to which the solar cell 2 is attached and the structure panel 1b to which the emissivity variable element 5 is attached. Thereby, excess heat generated in the solar cell 2 is transmitted through the heat pipes 6 a and 6 b and is radiated by the emissivity variable element 5. The heat pipes 6a and 6b embedded in the structure panels 1a and 1b may be a single heat pipe processed into an L shape.

  The solar cell 2 and the emissivity variable element 5 are directly attached to the structure panels 1a and 1b with an adhesive 6 having good thermal conductivity in order to enhance the thermal coupling with the structure panels 1a and 1b. Further, the structure panels 1a and 1b are made of a material having good heat conduction.

The emissivity variable element 5 is an element whose emissivity is actively changed according to temperature. Since the emissivity increases in a high temperature environment, the amount of heat dissipation increases, and in the low temperature environment, the emissivity decreases, so the amount of heat dissipation decreases.
When the temperature rises greatly when the condensed sunlight 4 is irradiated to the solar cell 2, the heat generated in the solar cell 2 is transferred from the solar cell 2 to the structure panels 1a and 1b, and the emissivity is variable. The element 5 radiates heat to the outside of the spacecraft. On the other hand, when sunlight 4 enters the shadow of a planet or the like and is not condensed, the flow of heat in the emissivity variable element 5 is shielded, making it difficult for heat to flow outside the spacecraft. Can be reduced.

  The embodiment disclosed this time is an example, and the present invention is not limited to this. The present invention is shown not by the above description but by the scope of the claims, and is intended to include all modifications within the meaning and scope equivalent to the scope of the claims.

A case where the present invention is applied to a Jupiter probe will be described. The Jupiter spacecraft is premised on the fact that it requires a generated power amount of about 300W in operation. In the vicinity of Jupiter, the sunlight intensity is about 50 W / m ² , so that it is necessary to set the area of the solar cell panel to 24 m ² in order to generate about 300 W of electric power by the conventional method. On the other hand, when the present invention is applied, the solar cell panel becomes unnecessary, and the mass of the solar cell panel per unit area is about 4 kg / m ^2, so that the mass of about 96 kg can be reduced.

Next, the temperature in the case of condensing on the solar cell attached to the structure panel using the communication antenna will be described. Assuming that two communication antennas are mounted on the spacecraft, the opening area of each communication antenna is about 12 m ² . Moreover, about 600 W of sunlight per surface is collected by the communication antenna and irradiated to the solar cell, and about 430 W of this is absorbed as heat into the spacecraft.

  When power of about 300 W is consumed inside the spacecraft and converted into heat, the amount of heat generated by the internal device per structure panel is about 75 W. When the conventional method is applied, the temperature of the structure panel to which the solar cell is attached is about 62 ° C. Since the allowable temperature on the high temperature side of internal equipment mounted on a general spacecraft is 55 to 65 ° C., there is a possibility of deviating from the allowable temperature of the internal equipment. On the other hand, when the exhaust heat system using the heat pipe and the emissivity variable element shown in the present invention is applied, the temperature of the internal device can be lowered to about 40 ° C.

  Furthermore, in the case where sunlight enters the shadow of a planet or the like and is not collected, if the variable emissivity element is not applied, about 100 W of heat is dissipated outside the spacecraft. On the other hand, when the present invention is applied, about 40 W of heat is radiated from the variable emissivity element to the outside of the spacecraft. For this reason, by applying the present invention, the electric power necessary for maintaining the temperature can be reduced by about 60 W.

DESCRIPTION OF SYMBOLS 1 Space organization panel (Structure panel), 2 Solar cell, 3 Communication antenna which added sunlight condensing function, 4 Sunlight, 5 Emissivity variable element, 6 Heat pipe, 7 Internal equipment

Claims (3)

A spacecraft equipped with a communication antenna,
The spacecraft which comprises the said spacecraft is equipped with the solar cell which light-receives the sunlight condensed with the said communication antenna, The spacecraft characterized by the above-mentioned. The spacecraft according to claim 1, wherein the opening surface of the communication antenna is subjected to a coating process, and the reflectance of the sunlight is increased by the coating process. The structure panel including the solar cell includes a heat pipe, and the heat pipe diffuses heat generated in the solar cell into the spacecraft. Spacecraft.

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