JP2013233906A - Spacecraft - Google Patents

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JP2013233906A
JP2013233906A JP2012109014A JP2012109014A JP2013233906A JP 2013233906 A JP2013233906 A JP 2013233906A JP 2012109014 A JP2012109014 A JP 2012109014A JP 2012109014 A JP2012109014 A JP 2012109014A JP 2013233906 A JP2013233906 A JP 2013233906A
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spacecraft
solar cell
sunlight
panel
communication antenna
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Tatsuya Kusajima
達也 草島
Shintaro Yamashita
慎太郎 山下
Toyoaki Funao
豊朗 舟生
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Mitsubishi Electric Corp
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Mitsubishi Electric Corp
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    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02EREDUCTION OF GREENHOUSE GAS [GHG] EMISSIONS, RELATED TO ENERGY GENERATION, TRANSMISSION OR DISTRIBUTION
    • Y02E10/00Energy generation through renewable energy sources
    • Y02E10/50Photovoltaic [PV] energy
    • Y02E10/52PV systems with concentrators

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

本発明は、太陽光を集光する機能を付加した通信アンテナを搭載した人工衛星等の宇宙機に関するものである。より詳しくは、宇宙機構体パネル外面に太陽電池を取り付けることにより宇宙機の小型・軽量化を図り、また、集光された太陽光の過剰な熱入力を効率的に宇宙機外部へ排熱することにより温度制御の効率化を図った宇宙機に関するものである。   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.

静止軌道あるいは低高度周回軌道上で運用されている宇宙機では、一般的に太陽電池を用いて発電が行われている。宇宙機の姿勢制御方式には、主にスピン安定方式及び三軸安定方式があり、これらの方式によって太陽電池の取り付け方式が異なる。   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.

スピン安定方式では、円筒状の宇宙機構体の外周面に、太陽電池を直接貼り付けた構成となっている。太陽光は円筒状の宇宙機構体側面から入射するため、貼り付けた太陽電池の半分程度にしか同時に照射されない。また、電力発生量は太陽電池を貼り付ける宇宙機構体の寸法に制限され、十分な電力が確保できないという問題点があった。   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.

これに対し、三軸安定方式では、直方体状の宇宙機構体パネルから太陽光の照射方向と対向する平面上に、太陽電池を多数に取り付けた太陽電池パネルを展開する構成となる。太陽電池パネルには、常に太陽と対向するように駆動機構が設けられている。これにより、常時、太陽電池パネル全面に太陽光を受光することが可能となり、十分な電力を確保することができる。この方式が、現在の宇宙機用発電システムの主流となっている。   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.

太陽電池の発電量は、設置面積及び太陽光強度等に比例する。太陽光強度は、太陽と宇宙機の位置関係によって決定されるため、宇宙機には必要な発電量に応じて、太陽電池パネル面積を調整している。   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.

地球を周回する宇宙機においては、宇宙機と太陽との距離が近く、太陽光強度が大きいため、太陽電池パネルのサイズは小さい。   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.

しかしながら、地球外惑星等を探査する場合、太陽との距離が遠くなるため、宇宙機は太陽光強度の弱い環境下におかれることになる。すなわち、地球を周回する宇宙機に比べ単位面積当たりの発電量は低下することになる。木星探査機を具体例として挙げると、木星近傍の太陽光強度は地球近傍の約4%となり、木星探査機の単位面積当たりの発電量は、地球を周回する宇宙機の約4%となる。その結果、地球を周回する宇宙機と同一電力量を供給するためには、太陽電池パネルの大きさを約25倍にする必要がある。   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.

巨大化する太陽電池パネルの問題に対し、特許文献1は、太陽電池パネルの周囲に反射鏡を取り付け、太陽電池パネルに太陽光を集光する技術を開示している。特許文献1の技術は、反射鏡に照射した太陽光が全て太陽電池パネルに集光される構成となっており、太陽電池パネルの面積を約半分程度に削減することを可能にしている。   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.

また、特許文献2では、半球状の反射鏡に照射した太陽光を反射鏡の焦点付近に配置した太陽電池に集光し発電を行う技術を開示している。この技術は宇宙機に対し適用することを想定した技術ではないが、この技術を活用することにより太陽電池の発電効率を高めることができ、太陽電池の数量を大幅に削減することが可能となる。   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. .

特開平11−245897号公報Japanese Patent Laid-Open No. 11-245897 特開2011−129847号公報JP 2011-129847 A

このように地球外惑星等を探査する場合、太陽との距離が遠く、太陽光強度が弱くなる。太陽光強度が弱い環境下において、従来方式を適用した場合、宇宙機の運用に必要な電力を得るためには、巨大な太陽電池パネルを取り付ける必要がある。巨大な太陽電池パネルを取り付けると、宇宙機の質量は増大するという課題があった。
質量が増大することにより、宇宙機の姿勢制御及び軌道制御を行う際に使用する推進剤の使用量が多くなるため、宇宙機の設計寿命が短縮化するという課題があった。
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.

このような課題に対し、従来、特許文献1及び特許文献2にあるように、太陽電池パネルを反射鏡等と組み合わせる技術が開示されている。これらの技術は、反射鏡等を使用しない場合において、必要となる太陽電池の面積と等価あるいはそれ以上の面積を有する反射鏡を取り付ける必要がある。すなわち、特許文献1および特許文献2の技術では、太陽電池の面積は削減できるが、発電用に大型の展開型構造物が必要なことに変わりはなく、上述の課題を解決するに至らない。
また、太陽電池で吸収された太陽光の一部は電力に変換されるが、残りの太陽光は熱に変換される。特許文献1及び特許文献2のように太陽光を集光した場合、単位面積当りの発電量は増加するが、太陽光による熱入力量も増加し、太陽光を集光した部分の温度が上昇する。太陽電池が高温化することにより、太陽電池の性能の低下、電気回路の故障等の別の問題が発生する。このように宇宙機において太陽光を集光し発電する場合、過剰な熱入力が別の課題として生じてしまう。
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.

本発明は係る課題を解決するためになされたものであり、宇宙機に搭載される太陽電池パネルを削減し、宇宙機の質量を軽量化できる太陽光集光通信アンテナを提供することを目的とする。   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.

この発明に係る宇宙機は、通信アンテナを備えた宇宙機であって、前記宇宙機を構成する構体パネルは、前記通信アンテナで集光された太陽光を受光する太陽電池を備える。   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.

この発明に係る太陽光集光通信アンテナによれば、従来と同じく通信用アンテナとして使用しつつ、太陽光を集光することができる。これにより、新たな構造物を追加することなく太陽電池パネルを削減し、宇宙機の質量を軽量化できる。   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.

本発明の実施の形態1に係る太陽光集光通信アンテナの構成を示す概略図である。It is the schematic which shows the structure of the sunlight condensing communication antenna which concerns on Embodiment 1 of this invention. 本発明の実施の形態1に係る太陽光集光通信アンテナの部分断面図である。It is a fragmentary sectional view of the sunlight condensing communication antenna which concerns on Embodiment 1 of this invention.

実施の形態1.
図1は、本発明の実施の形態1に係る人工衛星等の宇宙機の構成の概略図である。宇宙機構体パネル1a(以下、単に構体パネル1aという)には、太陽光反射率が大きくなる表面処理等を施した通信アンテナ3が取り付けられる。
通信アンテナ3は例えば地上局との間で通信を行う通信機能のほか、太陽光を反射し集光する機能を備える。通信アンテナ3は構体パネル1aから展開し、地球を指向するような配置をとる。
構体パネル1aには、通信アンテナ3の焦点となる位置に太陽電池2を取り付ける。通信アンテナ3の開口面に照射された太陽光4を反射し、構体パネル1aに取り付けた太陽電池2に集光させ発電を行う。
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.

通信アンテナ3には、電波及び太陽光4の両方を反射可能にするため、パラボラアンテナ、球面アンテナ等を用いる。アンテナ開口面には、太陽光反射率を高めるため、鏡面加工等の表面処理が施される。例えば、太陽光反射率を高めるためアンテナ開口面の表面にコーティング処理を行う。また、アンテナ開口面に電波を反射し、かつ太陽光反射率が大きい素材を取り付けたものであっても良い。   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.

本発明に係る実施の形態の構成を示す部分断面図を図2に示す。太陽電池2で生じた熱を宇宙機内部に拡散するため、構体パネル1aには熱輸送能力の高いヒートパイプ6aが埋め込まれる。また、宇宙機構体パネル1b(以下、構体パネル1b)には、内部機器7の発熱及び太陽電池2で生じた熱を放熱するため、放射率可変素子5が取り付けられる。なお、構体パネル1bにも、ヒートパイプ6bが埋め込まれる。   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.

太陽電池2が取り付けられる構体パネル1aと放射率可変素子5が取り付けられる構体パネル1bとの熱結合を強めるため、ヒートパイプ6aとヒートパイプ6bは連結される。これにより、太陽電池2で生じた余分な熱は、ヒートパイプ6a、6bを伝わり、放射率可変素子5にて放熱される。なお、構体パネル1a、1bに埋め込まれるヒートパイプ6a、6bは、1本のヒートパイプをL字型に加工したものであっても良い。   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.

太陽電池2及び放射率可変素子5は、構体パネル1a、1bとの熱結合を強めるため、熱伝導率の良い接着剤6で構体パネル1a、1bに直接貼り付けられる。また、構体パネル1a、1bは熱伝導の良い素材で構成される。   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.

放射率可変素子5は、温度に応じて能動的に放射率が変化する素子である。高温環境では放射率が大きくなるため放熱量が増加し、低温環境では放射率が小さくなるため放熱量が減少する。
集光された太陽光4が太陽電池2に照射されることによって温度が大きく上昇する場合、太陽電池2で生じた熱は、太陽電池2から構体パネル1a、1bに熱が伝わり、放射率可変素子5にて宇宙機外部へ放熱される。一方で、惑星等の影に入り太陽光4が集光されない場合、放射率可変素子5での熱の流れは遮蔽され、宇宙機外部に熱が流れにくい構成となり、温度保持に必要な電力を削減することができる。
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.

今回開示した実施の形態は例示であってこれに制限されるものではない。本発明は上述した内容ではなく、特許請求の範囲によって示され、特許請求の範囲と均等の意味及び範囲でのすべての変更が含まれることが意図される。   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.

木星探査機に本発明を適用した場合について説明する。木星探査機は、運用において約300Wの発生電力量を必要とすることを前提とする。木星近傍では太陽光強度が約50W/mであることから、従来方式で約300Wの電力を発電するためには、太陽電池パネルの面積を24mとする必要がある。これに対し本発明を適用した場合、太陽電池パネルは不要となり、単位面積当りの太陽電池パネルの質量は4kg/m程度であるから、約96kgの質量を削減することが可能となる。 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.

次に、通信アンテナを用いて、構体パネルに取り付けた太陽電池に集光した場合の温度について説明する。通信アンテナは宇宙機に2枚搭載することを前提とすると、1枚当りの通信アンテナの開口面積は約12mとなる。また、太陽電池には、1面当り約600Wの太陽光が通信アンテナで集光されて照射し、このうち約430Wが熱として宇宙機内部に吸収される。 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.

約300Wの電力を宇宙機内部で消費され熱に変換されると、構体パネル1枚当りの内部機器発熱量は約75Wとなる。従来の方式を適用した場合、太陽電池を取り付けた構体パネルの温度は約62℃となる。一般的な宇宙機に搭載される内部機器の高温側の許容温度は55〜65℃でありため、内部機器の許容温度を逸脱する可能性がある。これに対し、本発明に示すヒートパイプと放射率可変素子を用いた排熱システムを適用した場合、内部機器の温度を約40℃に温度を下げることが可能となる。   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.

さらに、惑星等の影に入り太陽光が集光されない場合においては、放射率可変素子を適用しない場合、約100Wの熱量が宇宙機外部に放熱される。一方、本発明を適用した場合、約40Wの熱量が放射率可変素子から宇宙機外部に放熱される。このことから、本発明を適用することにより、温度保持に必要な電力を約60W削減することができる。   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.

1 宇宙機構体パネル(構体パネル)、2 太陽電池、3 太陽光集光機能を付加した通信アンテナ、4 太陽光、5 放射率可変素子、6 ヒートパイプ、7 内部機器 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)

通信アンテナを備えた宇宙機であって、
前記宇宙機を構成する構体パネルは、前記通信アンテナで集光された太陽光を受光する太陽電池を備えることを特徴とする宇宙機。
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.
前記通信アンテナの開口面にはコーティング処理が施され、コーティング処理により前記太陽光の反射率が高くなることを特徴とする請求項1記載の宇宙機。 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. 前記太陽電池が備えられた前記構体パネルにはヒートパイプが備えられ、前記ヒートパイプは前記太陽電池で生じた熱を前記宇宙機内部に拡散させることを特徴とする請求項1、2いずれか記載の宇宙機。 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.
JP2012109014A 2012-05-11 2012-05-11 Spacecraft Pending JP2013233906A (en)

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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2017046428A (en) * 2015-08-25 2017-03-02 大和ハウス工業株式会社 Power interchange system
CN111338404A (en) * 2020-02-27 2020-06-26 北京空间飞行器总体设计部 Satellite power temperature control method
US11414220B2 (en) 2016-03-31 2022-08-16 Mitsubishi Electric Corporation Heat radiator using heat pipe panel

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2017046428A (en) * 2015-08-25 2017-03-02 大和ハウス工業株式会社 Power interchange system
US11414220B2 (en) 2016-03-31 2022-08-16 Mitsubishi Electric Corporation Heat radiator using heat pipe panel
CN111338404A (en) * 2020-02-27 2020-06-26 北京空间飞行器总体设计部 Satellite power temperature control method
CN111338404B (en) * 2020-02-27 2021-09-24 北京空间飞行器总体设计部 Satellite power temperature control method

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