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WO2018030368A1 - Sonde, procédé de fabrication d'élément de sonde et procédé de fabrication de sonde - Google Patents

Sonde, procédé de fabrication d'élément de sonde et procédé de fabrication de sonde Download PDF

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Publication number
WO2018030368A1
WO2018030368A1 PCT/JP2017/028682 JP2017028682W WO2018030368A1 WO 2018030368 A1 WO2018030368 A1 WO 2018030368A1 JP 2017028682 W JP2017028682 W JP 2017028682W WO 2018030368 A1 WO2018030368 A1 WO 2018030368A1
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WO
WIPO (PCT)
Prior art keywords
spacecraft
camera
housing
probe
plate
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Ceased
Application number
PCT/JP2017/028682
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English (en)
Japanese (ja)
Inventor
ジョン ウォーカー
敏郎 清水
利樹 田中
大輔 古友
裕 工藤
清菜 宮本
武史 袴田
貴裕 中村
大士 松倉
モハメド ラガブ
アブデルカデル ハウシン
ダミヤン ハイカル
チイホン ヤン
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Ispace Inc
Original Assignee
Ispace Inc
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Ispace Inc filed Critical Ispace Inc
Publication of WO2018030368A1 publication Critical patent/WO2018030368A1/fr
Anticipated expiration legal-status Critical
Ceased legal-status Critical Current

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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/16—Extraterrestrial cars
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01V—GEOPHYSICS; GRAVITATIONAL MEASUREMENTS; DETECTING MASSES OR OBJECTS; TAGS
    • G01V9/00—Prospecting or detecting by methods not provided for in groups G01V1/00 - G01V8/00

Definitions

  • the present disclosure relates to a probe, a method of manufacturing a probe component, and a probe manufacturing method.
  • Spacecraft used for lunar or planetary exploration activities are known.
  • a spacecraft there is a space exploration vehicle that can travel on the moon surface or on the planet (refer to Japanese Patent Application Laid-Open No. 2010-132261), and a US Mars rover is known.
  • a spacecraft is a travelable spacecraft, and can travel a wheel, a first camera arranged in a direction in which the spacecraft can travel, and the spacecraft.
  • a second camera arranged in a direction other than the direction, and the lens of the first camera and / or the second camera is oriented downward from the horizontal, and the first camera Wheels are included in the camera field of view and / or in the field of view of the second camera.
  • FIG. 27 is a cross-sectional view of the hub HB when cut along the DD cross section of FIG. 26. It is sectional drawing of the clamp HC when it cuts in CC section of FIG. It is a schematic diagram which shows the usage pattern of the heat insulation sheet which concerns on 2nd Embodiment.
  • This disclosure has been made in view of the above problems, and an object of the present disclosure is to provide a probe that can be searched even if it is downsized.
  • the spacecraft according to the first aspect of the present disclosure is a travelable spacecraft, and includes a wheel, a first camera arranged in a direction in which the spacecraft can travel, and the spacecraft traveling.
  • This configuration makes it possible to confirm whether or not stones are caught in the wheels.
  • the spacecraft according to the second aspect of the present disclosure is the spacecraft according to the first aspect, and the resolution of the first camera is higher than the resolution of the second camera.
  • This configuration allows the field of view in the direction of travel to be viewed with higher resolution, so that obstacles in the direction of travel can be easily found.
  • a spacecraft according to a third aspect of the present disclosure is a spacecraft according to the first or second aspect, and includes a plurality of processors, and the spacecraft is capable of traveling both in the front and rear directions.
  • a camera it has a front camera arranged toward the front and a rear camera arranged toward the rear, and the front camera and the rear camera are respectively connected to separate processors.
  • the spacecraft can be moved either forward or backward while viewing the image of either the front camera or the rear camera.
  • a probe according to a fourth aspect of the present disclosure is the probe according to any one of the first to third aspects, and is connected to the first camera or the second camera via a serial or parallel interface.
  • a camera controller, and a communication controller for communicating the camera controller requests and obtains video data from the camera at a predetermined frame rate, and compresses the obtained video data by hardware encoding.
  • the communication controller transmits the compressed data.
  • moving image data captured by the probe camera can be transferred to the ground station, and the operator operating the probe can view the moving image data on the earth.
  • the first camera or the second camera need not always be turned on, and only needs to operate when moving image data is requested, so that power consumption can be suppressed.
  • a probe according to a fifth aspect of the present disclosure is the probe according to any one of the first to fourth aspects, and includes a housing, and the housing includes a substrate, a Teflon (registered trademark) layer, A quartz glass layer, and a metal film provided between the substrate and the Teflon (registered trademark) layer or the quartz glass layer.
  • a probe according to a sixth aspect of the present disclosure is the probe according to the fifth aspect, and an indium tin oxide layer is provided on the Teflon (registered trademark) layer or the quartz glass layer. .
  • a probe according to a seventh aspect of the present disclosure is a probe according to any one of the first to sixth aspects, and includes a housing and an electronic device, and the housing includes a side plate, The top plate to which the electronic device is fixed is provided, and a heat insulating material is provided between the side plate and the top plate.
  • the probe according to the eighth aspect of the present disclosure is the probe according to the seventh aspect, and the electronic device is provided on the back of the top board.
  • a spacecraft according to a ninth aspect of the present disclosure is the spacecraft according to any one of the first to eighth aspects, wherein the spacecraft is exposed to the top plate in a state of being exposed from the top surface of the housing and the housing.
  • the heat generated from the battery can be released from the top plate of the housing to the outer space to suppress the temperature rise of the battery.
  • a probe according to a tenth aspect of the present disclosure is the probe according to any one of the first to ninth aspects, and includes a housing, and the front plate and / or the rear plate of the housing is formed from a bottom plate. It leans to the inside of the spacecraft over the top plate.
  • This configuration can reduce the rate at which sunlight reflected on the lunar surface hits the front plate and the rear plate, so that the temperature rise of the spacecraft can be suppressed.
  • a probe according to an eleventh aspect of the present disclosure is the probe according to any one of the first to tenth aspects, with respect to a contour line between a casing and a bottom surface of a side plate of the casing.
  • the solar cells are arranged obliquely with respect to the perpendicular.
  • This configuration makes it possible to increase the exclusive area ratio of the outer surface of the solar cell housing, and to arrange a large number of solar cells within a limited area.
  • a probe according to a twelfth aspect of the present disclosure is the probe according to the eleventh aspect, further comprising a charge / discharge circuit to which power generated by the solar cell is supplied, and penetrating the casing. A hole is provided, and wiring from the solar cell is connected to a charge / discharge circuit in the casing through a through hole provided in the casing.
  • This configuration eliminates the need to provide a space for fixing the wiring on the outer surface of the housing, so that many solar cells can be arranged within a limited area.
  • a probe according to a thirteenth aspect of the present disclosure is the probe according to any one of the first to twelfth aspects, further including a solar cell disposed on the housing, wherein the solar cell is disposed.
  • the surface of the case is tilted to the inside of the spacecraft from the bottom plate to the top plate.
  • This configuration makes it possible to efficiently receive light from the sun, thus increasing the amount of power generation.
  • the probe according to the fourteenth aspect of the present disclosure is the probe according to the thirteenth aspect, and the inclination of the surface of the housing in which the solar cell is disposed is the latitude at which the probe is to be disposed. It is decided according to.
  • This configuration sets the solar cell inclination according to the maximum elevation angle of the sun, so that the amount of power generation can be increased.
  • a probe according to a fifteenth aspect of the present disclosure is the probe according to any one of the first to fourteenth aspects, and includes a motor provided on the wheel and a cone-shaped convex portion in the vicinity of the center.
  • the convex portion is fitted in the second hole in a state where the back surface of the clamp and the hub are opposed to each other.
  • a spacecraft according to a sixteenth aspect of the present disclosure is a spacecraft, is folded and stored, and includes a heat insulating sheet that can be expanded from a folded state. It is configured to cover the outside of the spacecraft.
  • a spacecraft includes an interface connected to a payload or an accessory, and when the interface is connected to the payload or the accessory, the interface converts the voltage to the payload or the accessory to generate power.
  • This configuration provides power to the payload or accessory and allows the spacecraft to exchange electrical signals with the payload or accessory.
  • the method for manufacturing a probe component according to the eighteenth aspect of the present disclosure includes a step of manufacturing the probe component using a 3D printer disposed on a celestial body other than the earth.
  • the probe component when a probe component fails or is damaged, the probe component can be manufactured on a celestial body other than the earth with a 3D printer, and the manufactured component can be replaced with the failed component.
  • a method of manufacturing a probe component according to a nineteenth aspect of the present disclosure is a method of manufacturing a probe component according to the eighteenth aspect, the step of melting the probe component that has failed or is damaged, In the manufacturing step, spacecraft parts are manufactured by a 3D printer using the material after melting as a raw material.
  • This configuration allows the spacecraft parts to be manufactured by reusing the spacecraft parts when the spacecraft parts break down or are damaged.
  • a method for manufacturing a probe component according to a twentieth aspect of the present disclosure is a method for manufacturing a probe component according to the eighteenth aspect, the step of collecting natural resources in a celestial body other than the earth, and the manufacture In this step, the parts of the probe are manufactured by a 3D printer using the collected natural resources as raw materials.
  • This configuration makes it possible to manufacture spacecraft parts at a lower cost.
  • a probe manufacturing method is a probe manufacturing method for manufacturing a probe in a celestial body other than the earth, in which a natural resource is collected from a celestial body other than the earth, or a fault or damage is detected.
  • This configuration makes it possible to manufacture a spacecraft at low cost, or to replace a damaged or failed spacecraft component at low cost. Alternatively, even if the parts of the probe are broken or damaged, the probe can be regenerated by regenerating and replacing the parts of the probe.
  • a spacecraft is based on a solar panel, a drive mechanism that changes the inclination of the solar panel from a horizontal plane, and the time, the position of the sun, or the amount of power generated by the solar panel. And a controller for controlling the drive mechanism so as to change the inclination with respect to the horizontal plane of the solar panel.
  • the inclination with respect to the horizontal plane of the solar panel can be changed according to the irradiation angle of sunlight, and the amount of light hitting the solar panel can be increased. Can be increased.
  • the spacecraft according to the twenty-third aspect of the present disclosure includes a reflector that reflects light, and a direction of the reflector so that sunlight reflected by the reflector is applied to a solar panel of an object. And / or a controller that controls to change the angle.
  • the reflected light can be applied to the solar panel of the object, so that the solar panel of the object can generate power.
  • a spacecraft includes a camera, an injection mechanism that can inject the camera, and a controller that controls the camera and the injection mechanism. Control is performed so that the camera takes a picture at the landing point, and an image obtained by the photography is acquired from the camera.
  • a spacecraft includes a casing, a switching mechanism that switches between an open state that releases heat in the casing and a blocking state that blocks heat in the casing, and an exterior of the casing Alternatively, a temperature sensor that measures the internal temperature and a processor that controls the switching mechanism to switch between an open state and a shut-off state according to the temperature measured by the temperature sensor.
  • a spacecraft includes a housing, a support post coupled to the housing, two pairs of legs connected to the support at one end, and the other ends of the legs. And the two pairs of legs are configured to be rotatable with respect to the column so that the angle between the two pairs of legs is variable.
  • This configuration lowers the center of gravity, so if the vibration during transportation is large, it can reduce the shake of the probe due to the vibration and suppress the damage caused by the shake. Further, when the spacecraft or the lander is loaded with the spacecraft as a payload, the height can be reduced by folding the legs, so that the space occupied by the spacecraft when loaded is reduced.
  • the spacecraft according to the twenty-seventh aspect of the present disclosure includes a housing including graphene or graphene fiber as a material.
  • This configuration can improve the heat insulation of the housing.
  • FIG. 1 is a schematic diagram showing an outline of an exploration system according to the present embodiment.
  • the exploration system S includes an exploration device (also referred to as a rover) R for exploring the lunar surface LS, and a landing ship (also referred to as a lander) L for transporting the exploration device to the moon MN.
  • a ground station E provided on the earth ET.
  • the spacecraft R according to the present embodiment is an unmanned spacecraft as an example, and can travel on the moon surface.
  • the probe R can communicate with the lander L.
  • the landing ship L can communicate with the ground station. Thereby, the spacecraft E can be controlled from the ground station E.
  • FIG. 2 is a perspective view showing an outline of the spacecraft according to the present embodiment.
  • the spacecraft R according to the present embodiment includes a housing HS, shafts LSF and RSF provided in the housing HS, wheels FW1 and RW1 connected to a shaft RSF (not shown), Wheels FW2 and RW2 connected to the shaft LSF are provided.
  • the spacecraft R includes a distance sensor DS provided on the front surface of the casing, and a first antenna AT1 and a second antenna AT2 provided on the top plate of the casing HS.
  • the distance sensor DS measures a distance from an object on the moon (for example, an obstacle such as a rock).
  • the probe R according to the present embodiment has no difference in driving mechanism between the case of traveling in the direction in which the distance sensor DS is provided and the case of traveling in the opposite direction.
  • the direction where the distance sensor DS is provided is assumed to be the front, and the opposite direction is assumed to be the rear.
  • the spacecraft R includes a front camera FC, a rear camera BC, a right side camera RC, and a left side camera LC.
  • the front camera FC, the rear camera BC, the right-side camera RC, and the left-side camera LC have a lens and an imaging unit that captures an object using light incident from the lens. As shown in FIGS. 13A and 13C described later, the probe R can move back and forth, but cannot move left and right.
  • the front camera FC and the rear camera BC are an example of a first camera arranged in a direction in which the spacecraft R can travel.
  • the right side camera RC and the left side camera LC are an example of a second camera arranged in a direction other than the direction in which the spacecraft can travel.
  • FIG. 3 is a front view of the spacecraft according to the present embodiment as viewed from the front. As shown in FIG. 3, the wheel FW1 is connected to the shaft RSF, and the wheel FW2 connected to the shaft LSF is connected.
  • FIG. 4 is a side view of the spacecraft according to the present embodiment as viewed from the left side.
  • the directions of the lenses LF and LB of the front camera FC and the rear camera BC, which are the first cameras, are directed downward from the horizontal. Thereby, the obstacle on the moon surface in the running direction can be visualized.
  • FIG. 5 is a top view of the spacecraft according to the present embodiment as viewed from above.
  • the solar cells M7-1 to M7-4, M8 to M8-5, M9-1 to M9-5, M10-1 to M10- are also applied to the right side plate RP on the right side of the housing HS. 4, M11-1 to M11-5, and M12-1 to M12-5 are provided.
  • the solar cell when assuming a perpendicular along the side plate of the housing HS with respect to the contour line between the bottom surface of the side plate of the housing HS, the solar cell is It is arranged diagonally.
  • casing HS outer surface of a solar cell can be made high, and many solar cells can be arrange
  • the housing HS has a top plate TP, a front plate FP, a rear plate BP, a right side plate RP, a left side plate LP, and a bottom plate DP (not shown).
  • the right side plate RP or the left side plate LP may be collectively referred to as a side plate.
  • the plates PL1, PL2, PL3, and PL4 are fixed to the top plate TP in a state where they are exposed from the top plate TP of the housing HS. That is, the surface is exposed to the outside by connecting to the top plate TP of the housing HS.
  • FIG. 6 is a schematic diagram of the AA cross section of FIG.
  • the spacecraft according to the present embodiment searches outside the equator of the moon. That is, it is assumed that sunlight is incident on the spacecraft R at an angle. For this reason, as shown in FIG. 6, the electronic device is provided in the back of the top plate, and the electronic device is being fixed to the top plate TP. Thereby, since reflected light when sunlight reflects on the ground such as the lunar surface LS does not hit the top plate TP, the electronic device is provided on the back of the top plate to prevent the temperature of the electronic device from rising. can do.
  • a battery board BB on which a battery which is one of electronic devices is mounted is fixed to the back surface of the plate PL1 via support columns P1-1 and P1-2.
  • the plate PL1 has a convex cross section, and is fitted into an opening provided in the top plate TP of the housing HS.
  • the battery which is one of the electronic devices, is fixed to the top plate TP.
  • an adhesive material (gel) GL1 having high thermal conductivity is sandwiched between the battery board BB and the back surface of the top plate TP. Thereby, the heat generated in the battery can be efficiently transmitted to the plate PL1, and the heat dissipation effect can be improved.
  • a power supply board PUB on which a power supply controller, which is one of electronic devices, is mounted is fixed to the back surface of the plate PL2 via support columns P2-1 and P2-2.
  • the plate PL2 has a convex cross section, and is fitted in an opening provided in the top plate TP of the housing HS.
  • the power supply controller that is one of the electronic devices is fixed to the top plate TP.
  • an adhesive material (gel) GL2 having high thermal conductivity is sandwiched between the power supply board PUB and the back surface of the top plate TP. Thereby, the heat generated by the power supply controller can be efficiently transmitted to the plate PL2, and the heat dissipation effect can be improved.
  • a motor board MCB on which a motor controller which is one of electronic devices is mounted is fixed to the back surface of the plate PL3 via support columns P3-1 and P3-2.
  • the plate PL3 has a convex cross section, and is fitted into an opening provided in the top plate TP of the housing HS.
  • the motor controller which is one of the electronic devices is fixed to the top plate TP.
  • an adhesive (gel) GL3 having a high thermal conductivity is sandwiched between the motor board MCB and the back surface of the top plate TP. Thereby, the heat generated by the motor controller can be efficiently transmitted to the plate PL3, and the heat dissipation effect can be improved.
  • a camera board CB on which a camera controller which is one of electronic devices is mounted is fixed to the back surface of the plate PL4 via support columns P4-1 and P4-2.
  • the plate PL4 has a convex cross section, and is fitted into an opening provided in the top plate TP of the housing HS.
  • the camera controller which is one of the electronic devices is fixed to the top plate TP.
  • an adhesive (gel) GL4 having a high thermal conductivity is sandwiched between the camera board CB and the back surface of the top plate TP. Thereby, the heat generated by the camera controller can be efficiently transmitted to the plate PL4, and the heat dissipation effect can be improved.
  • a communication board RB on which a communication controller that is one of electronic devices is mounted is fixed to the back surface of the plate PL5 via support columns P5-1 and P5-2.
  • the plate PL5 has a convex cross section, and is fitted into an opening provided in the rear plate BP of the housing HS.
  • the communication controller which is one of the electronic devices is fixed to the rear plate BP.
  • an adhesive (gel) GL4 having a high thermal conductivity is sandwiched between the camera board CB and the back surface of the top plate TP. Thereby, the heat generated by the camera controller can be efficiently transmitted to the plate PL4, and the heat dissipation effect can be improved.
  • FIG. 7 is a perspective view showing the structure of the plate PL4.
  • a camera board CB on which a camera controller is mounted and a plate PL4 are connected via four columns P4-1 to P4-4.
  • four holes HE1 to HE4 for fixing the plate PL4 to the top plate TP with screws are provided.
  • assembly can be simplified by packaging.
  • FIG. 8 is a table showing an example of the heat radiation amount for each plate and the paint color of the exposed surface (surface) of the plate.
  • the reflectance and absorption rate of light differ depending on the color. Therefore, the color of the exposed surface of the plate on which the electronic device is mounted is set so that the greater the heat dissipation amount of the electronic device, the higher the light reflectance and the lower the absorption rate.
  • the plates PL1 and PL4 to which the heat radiation source having a large heat radiation amount is fixed have, for example, a white paint color on the exposed surface (surface).
  • White has high light reflectivity and low absorptance, so it can suppress the temperature rise of the plates PL1 and PL4 even when exposed to sunlight, and can suppress the temperature rise of the battery and camera controller with large heat dissipation. it can.
  • the plates PL2 and PL3 to which the heat radiation source having a small heat radiation amount is fixed have, for example, a black painted color on the exposed surface (surface). Black has low light reflectivity and high absorption, so when sunlight is applied, it promotes the temperature rise of the plates PL2 and PL3, and makes the temperature of the battery and camera controller with a small amount of heat dissipation moderate. Can do.
  • the front plate FP and the rear plate BP of the housing HS are inclined inward of the spacecraft R from the bottom plate DP to the top plate TP.
  • the ratio which the sunlight reflected on the moon surface hits the front board FP and the back board BP can be reduced, the temperature rise of the spacecraft R can be suppressed.
  • FIG. 9 is a schematic view of the BB cross section of FIG.
  • the right side plate RP and the left side plate LP of the housing HS are inclined inward of the spacecraft R from the bottom plate DP to the top plate TP.
  • the ratio which the sunlight reflected on the lunar surface hits the right side board RP and the left side board LP can be reduced, the temperature rise of the spacecraft R can be suppressed.
  • a heat insulating material HI-1 is provided between the right side plate RP and the top plate TP, and the heat insulating material HI-1 is fixed to the top plate TP with bolts B1.
  • a heat insulating material HI-2 is provided between the left side plate LP and the top plate TP, and the heat insulating material HI-2 is fixed to the top plate TP with bolts B2.
  • the heat insulating materials HI-1 and HI-2 according to the present embodiment are, for example, engineering plastics, for example, ULTEM (registered trademark) of amorphous thermoplastic polyetherimide (PEI) resin.
  • FIG. 10 is a schematic diagram showing the visual field range of the camera in the horizontal direction.
  • the front view range FHV is the view range of the front camera FC
  • the rear view range BHV is the view range of the rear camera BC
  • the right view range RHV is the view range of the right camera RC
  • the left view range LHV is the left camera.
  • LC viewing range As shown in FIG. 10, the front visual field range FHV and the right side visual field range RHV partially overlap, and the front visual field range FHV and the left side visual field range LHV partially overlap.
  • the rear visual field range BHV and the right side visual field range RHV partially overlap, and the rear visual field range BHV and the left side visual field range LHV partially overlap. Thereby, 360 degrees around can be seen in the horizontal direction.
  • FIG. 11 is a schematic diagram showing the field of view of the camera in the AA section of FIG.
  • the front vertical visual field range FVV is a vertical visual field range of the front camera FC
  • the rear vertical visual field range is a vertical visual field range of the rear camera BC.
  • the front vertical visual field range FVV includes wheels FW1 and FW2.
  • wheels RW1 and RW2 are included in the rear vertical visual field range BVV.
  • both the front camera FC and the rear camera BC are in the field of view above the horizontal line L1.
  • FIG. 12 is a schematic diagram showing the field of view of the camera in the BB cross section of FIG.
  • the right vertical visual field range RVV is a vertical visual field range of the right-side camera RC
  • the left vertical visual field range is a vertical visual field range of the left-side camera LC.
  • the right vertical visual field range RVV includes wheels FW1 and RW1. Thereby, it can be confirmed whether or not stones are sandwiched between the wheels FW1 and RW1.
  • the left vertical visual field range LVV includes wheels FW2 and RW2. Thereby, it can be confirmed whether or not stones are sandwiched between the wheels FW2 and RW2.
  • both the right side camera RC and the left side camera LC are in the field of view above the horizontal line L2.
  • the resolution of the first camera such as the front camera FC and the rear camera BC is higher than the resolution of the second camera such as the right side camera RC and the left side camera LC. That is, the resolution of the camera in the traveling method (the method in which the wheel advances) is higher than the resolution of the side camera. Thereby, since the visual field range in the traveling direction can be seen with higher resolution, an obstacle or the like in the traveling direction can be easily found.
  • FIG. 13A is a schematic diagram illustrating a first movement mode of the spacecraft R.
  • FIG. 13B is a schematic diagram illustrating a second movement mode of the spacecraft R.
  • FIG. 13C is a schematic diagram showing directions in which the spacecraft R cannot move.
  • FIG. 13D is a schematic diagram illustrating a third movement mode of the spacecraft R. As shown to FIG. 13A, it can move to back and front, and as shown to FIG. 13B, it can turn on the spot. However, as shown in FIG.
  • FIG. 14 is a schematic block diagram showing the configuration of the spacecraft R according to the present embodiment.
  • the probe R includes a battery BAT and a power supply controller that controls the battery BAT.
  • the spacecraft R includes a motor MT, a gear box GB, and a motor controller MC that controls the motor MT and the gear box GB.
  • the spacecraft R is connected to the front camera FC, the right side camera RC, the first camera controller CMC1 that controls the front camera FC and the right side camera RC, and the first camera controller CMC1 via wiring.
  • 1 communication controller CC1 and 1st antenna AT1 connected to 1st communication controller CC1 are provided.
  • the first camera controller CMC1 includes a first processor PC1.
  • the spacecraft R is connected to the rear camera BC, the left camera LC, the second camera controller CMC2 for controlling the rear camera BC and the left camera LC, and the second camera controller CMC2 via wiring.
  • 2 communication controller CC2 and 2nd antenna AT2 connected to 2nd communication controller CC2.
  • the second camera controller CMC2 includes a second processor PC2.
  • the front camera FC is connected to the first processor PC1
  • the rear camera BC is connected to the second processor PC2. That is, the front camera and the rear camera are connected to separate processors.
  • the other processor can operate. Therefore, either the front camera FC or the rear camera BC can be operated. Can be transferred to the ground station E. Therefore, the spacecraft R can be moved either forward or backward while viewing the image of either the front camera FC or the rear camera BC.
  • FIG. 15 is a schematic diagram showing a hardware configuration of the first camera controller CMC1.
  • the front camera FC is connected to an A / D converter AD1
  • the A / D converter AD1 is connected to the serial interface SI1 of the first camera controller CMC1 via a flat cable.
  • the serial interface SI1 is an interface compliant with, for example, MIPI (Mobile Industry Processor Interface) standard.
  • the right side camera RC is connected to the A / D converter AD2, and the A / D converter AD2 is connected to the parallel interface PI1 of the first camera controller CMC1 via a flat cable.
  • the first camera controller CMC1 is connected to the front camera FC or the right camera RC via a serial or parallel interface.
  • the first camera controller CMC1 requests and acquires moving image data from the front camera FC and the right side camera RC at a predetermined frame rate, and compresses the acquired moving image data by hardware encoding.
  • the compressed data is transferred to the first communication controller CC1.
  • the first communication controller CC1 transmits the compressed data from the first antenna AT1 to the landing ship (lander) L. Thereafter, the compressed data is transferred from the lander L to the ground station E.
  • the moving image data photographed by the front camera FC and the right-side camera RC of the spacecraft R can be transferred to the ground station, and the operator who operates the spacecraft R can view the moving image data on the earth.
  • the operator who operates the spacecraft R can see the image of the moon on the earth.
  • the front camera FC and the right-side camera RC do not need to be always turned on like the USB camera, and need only operate when requesting moving image data, thereby reducing power consumption. Can do.
  • FIG. 16 is a schematic diagram showing a hardware configuration of the second camera controller CMC2.
  • the rear camera BC is connected to an A / D converter AD3, and the A / D converter AD3 is connected to the serial interface SI2 of the second camera controller CMC2 via a flat cable.
  • the serial interface SI2 is an interface compliant with, for example, MIPI (Mobile Industry Processor Interface) standard.
  • the left side camera LC is connected to the A / D converter AD4, and the A / D converter AD4 is connected to the parallel interface PI2 of the second camera controller CMC2 via a flat cable.
  • the second camera controller CMC2 is connected to the rear camera BC or the left camera LC via a serial or parallel interface.
  • the second camera controller CMC2 requests and acquires moving image data from the rear camera BC and the left camera LC at a predetermined frame rate, and compresses the acquired moving image data by hardware encoding.
  • the compressed data is transferred to the second communication controller CC2.
  • the second communication controller CC2 transmits the compressed data from the second antenna AT2 to the landing ship (lander) L. Thereafter, the compressed data is transferred from the lander L to the ground station E.
  • FIG. 17 is a schematic diagram illustrating an outline of a cross section of the housing HS.
  • the housing HS includes a substrate 1, a metal film (here, a silver film) 2 deposited on the substrate 1, and a Teflon ( Registered trademark) layer 3.
  • the substrate is, for example, carbon fiber reinforced plastic (Carbon Fiber Reinforced Plastics: hereinafter referred to as CFRP).
  • CFRP Carbon Fiber Reinforced Plastics
  • the housing HS has an indium tin oxide (hereinafter referred to as ITO) layer 4 provided on the Teflon (registered trademark) layer 3.
  • ITO is a transparent conductive film.
  • FIG. 18 is a flowchart illustrating an example of a process flow relating to coating of the casing on the front plate FP, the rear plate BP, the top plate TP, and the bottom plate DP.
  • Step S101 First, the CFRP plate is processed.
  • a CFRP plate is cut out to a predetermined size, and an opening for fitting a distance sensor is made.
  • CFRP is cut out to a predetermined size, and an opening for fitting the plate PL5 is made.
  • CFRP is cut out to a predetermined size, and openings for fitting the plates PL1 to PL4 are made.
  • CFRP is cut out to a predetermined size.
  • Step S102 silver is deposited on the CFRP plate in a vacuum.
  • Step S103 Teflon (registered trademark) powder is sprayed on the silver deposition surface.
  • Teflon (registered trademark) particles on the beads adhere to the silver deposition surface.
  • Step S104 Next, the temperature is raised and Teflon (registered trademark) is melted and baked. Thereby, the particles of Teflon (registered trademark) are melted and connected to each other, and the surface of the Teflon (registered trademark) layer becomes flat.
  • Step S105 ITO is deposited in vacuum. Thereby, the front plate FP, the rear plate BP, the top plate TP, and the bottom plate DP are obtained.
  • FIG. 19 is a schematic diagram showing the left side plate LP before coating.
  • the left side plate LP is provided with through holes H1 to H6 for drawing wirings connected to the solar cell into the housing HS. ing.
  • casing HS is similarly provided in the left side plate LP.
  • FIG. 20 is a diagram illustrating an example of the configuration of the power controller PU.
  • the solar cells are connected in series for each column by wiring, and the wiring is drawn into the housing HS from, for example, the through holes H1 to H6 shown in FIG. Connected.
  • each wiring from the solar cell is connected to the anodes of the corresponding diodes D1 to D12 and rectified.
  • the cathodes of the diodes D1 to D12 are respectively connected to the charge / discharge circuit CDC, and the current rectified by the diodes D1 to D12 is input to the charge / discharge circuit CDC.
  • the wiring from the solar cell is connected to the charge / discharge circuit CDC in the housing HS through the through holes H1 to H6 provided in the housing HS.
  • the charge / discharge circuit CDC charges the battery BAT using the input current.
  • the charge / discharge circuit CDC supplies power to other electronic devices using the power of the battery BAT.
  • FIG. 21 is a flowchart showing an example of a process of creating the right side plate RP or the left side plate LP which is a side plate.
  • Step S201 First, CFRP is cut into a predetermined size, and a through hole is opened in the CFRP plate.
  • Step S202 Next, the through hole provided in the CFRP plate is masked. Thereby, it can avoid that a through-hole is obstruct
  • Step S203 silver is deposited on the CFRP plate in a vacuum.
  • Step S204 Teflon (registered trademark) powder is sprayed onto the silver deposition surface.
  • Teflon (registered trademark) particles on the beads adhere to the silver deposition surface.
  • Step S205 Next, the temperature is raised and Teflon (registered trademark) is melted and baked. Thereby, the particles of Teflon (registered trademark) are melted and connected to each other, and the surface of the Teflon (registered trademark) layer becomes flat.
  • Step S206 ITO is deposited in vacuum.
  • Step S207 Next, the masking attached in Step S202 is taken.
  • Step S208 a heat-resistant and cold-resistant polyimide film is stuck on the ITO.
  • the polyimide film is, for example, Kapton (registered trademark).
  • Step S209 the solar cell is fixed on the polyimide film. Thereby, the right side plate RP or the left side plate LP which is a side plate is obtained.
  • FIG. 22 is a schematic diagram of a front view of the probe R at a predetermined latitude where the probe R is arranged.
  • the surface of the housing on which the solar cells are arranged (in this embodiment, the right side surface and the left side surface as an example) is inclined inward of the spacecraft R from the bottom plate DP to the top plate TP. ing. Thereby, since the light from the sun can be received efficiently, the power generation amount can be increased.
  • the right side surface and the left side surface which are the surfaces of the housing where the solar cells are disposed, are inclined at an angle at which the power generation capacity from when the sun rises to when it sinks is maximized at the latitude where the spacecraft R is to be disposed. Yes. Specifically, as shown in FIG. 22, when the maximum elevation angle of the sun at the latitude where the spacecraft R is to be arranged is ⁇ 1 degree, the angle at which the power generation capacity from when the sun rises until it sinks is maximum is ⁇ 1. Therefore, the inclination of the right side surface and the left side surface from the horizontal plane is set to ⁇ 1 degree.
  • FIG. 23 is a schematic diagram of a front view of the spacecraft R when the latitude where the spacecraft R is to be arranged is higher than that in FIG.
  • the maximum elevation angle ⁇ 2 of the sun at the latitude where the spacecraft R is to be arranged becomes smaller than ⁇ 1.
  • the angle at which the power generation capacity from when the sun rises until it sinks becomes ⁇ 2, which is larger than ⁇ 1, so the inclination of the right side surface and the left side surface from the horizontal plane is set to ⁇ 2 degrees.
  • FIG. 24 is a schematic diagram of a front view of the spacecraft R when the latitude where the spacecraft R is to be arranged is lower than in the case of FIG.
  • the maximum elevation angle ⁇ 3 of the sun at the latitude where the spacecraft R is to be arranged becomes larger than ⁇ 1.
  • the angle at which the power generation capacity from when the sun rises until it sinks becomes ⁇ 3 smaller than ⁇ 1, so the inclination of the right side surface and the left side surface from the horizontal plane is set to ⁇ 3 degrees.
  • the inclination of the surface (here, the right side surface and the left side surface) of the casing on which the solar cell is disposed is determined according to the latitude at which the spacecraft R is to be disposed. Specifically, the higher the latitude at which the spacecraft R is to be arranged, the smaller the maximum elevation angle of the sun, so the inclination of the right side RP and the left side from the horizontal plane increases. Thereby, since the inclination of the solar cell is set according to the maximum elevation angle of the sun, the amount of power generation can be increased.
  • FIG. 25 is an exploded perspective view of the wheel FW2.
  • the wheel FW2 includes a motor MT, a motor slave SV, a bearing BR1, a bearing spacer BS, a bearing BR2, a motor housing MH, a bearing hold plate BHP, a clamp HC, a hub HB, and a wheel WL2.
  • the motor MT is inserted into the motor slave SV, the rotation shaft of the motor MT is inserted into the first hole HL in FIG. 26 of the hub HB, and the rotation shaft of the motor MT is clamped to the hub HB by the clamp HC.
  • FIG. 26 is a front view of the hub HB as seen from the direction of the arrow A1 in FIG. As shown in FIG. 26, it has the 1st hole HL and the notch CO connected to the said 1st hole HL.
  • FIG. 27 is a cross-sectional view of the hub HB when cut along the DD cross section of FIG. As shown in FIG. 27, the hub HB has a cone-shaped projection PJ near the center. As described above, the hub HB has the cone-shaped convex portion PJ near the center, and communicates with the first hole HL in which the rotation shaft of the motor MT is fitted in the convex portion PJ and the first hole HL. Has cutout CO.
  • FIG. 28 is a cross-sectional view of the clamp HC taken along the CC cross section of FIG.
  • the rotating shaft of the motor MT is inserted from the front surface FS side, and the hub HB is on the back surface RS side.
  • the clamp HC has a second hole HL2 whose diameter gradually decreases from the rear surface RS toward the front surface FS.
  • the convex portion PJ is fitted in the second hole HL2 with the back surface RS of the clamp HC and the hub HB facing each other.
  • the notch CO is narrowed
  • the contour around the first hole HL of the hub HB is narrowed
  • the rotation shaft of the motor MT is strongly restrained. .
  • the motor MT is fixed, if an excessive force is applied to the rotating shaft of the motor MT in the rotating shaft direction, the motor MT suddenly stops rotating.
  • the electronic device is not fixed to the top plate, but the electronic device may be disposed on the back surface of the front plate FP or the rear plate BP, or the electronic device may be disposed on the back surface of the bottom plate DP.
  • an electronic device can be provided on the back side of the surface where sunlight does not enter, and an increase in temperature of the electronic device can be prevented.
  • the heat generated from the electronic device can be released from the top plate TP of the housing HS to the outer space to suppress the temperature rise of the electronic device.
  • both the front plate FP and the rear plate BP of the housing are inclined to the inside of the spacecraft R from the bottom plate DP to the top plate TP. Good.
  • both the right side plate RP and the left side plate LP of the housing are inclined inward of the spacecraft R from the bottom plate DP to the top plate TP, but only one of them is inclined. Also good.
  • the wheel FW1 is included in both the field of view of the front camera FC that is one of the first cameras and the field of view of the right-side camera RC that is one of the second cameras.
  • the present invention is not limited to this, and the wheel FW1 may be included only in one field of view.
  • the other wheels FW2, RW1, RW2. In this way, it is only necessary to see the wheels with at least one of the cameras. Thereby, it can be confirmed whether the wheel is not clogged with stones.
  • FIG. 29 is a schematic diagram illustrating a usage pattern of the heat insulating sheet according to the second embodiment.
  • the spacecraft RV1 is folded and stored, and includes a heat insulating sheet TT that can be expanded from the folded state.
  • the heat insulating sheet TT is configured to cover the outside of the spacecraft RV1 when it is spread. With this configuration, sunlight is reflected by the heat insulating sheet and has a heat insulating property, so that the temperature change of the spacecraft can be reduced.
  • the heat insulating sheet TT is preferably self-supporting.
  • the spacecraft RV1 may have a gas discharge mechanism that discharges gas into the heat insulation sheet TT and a controller that controls the gas discharge mechanism in a state where the heat insulation sheet TT is folded. Accordingly, the controller may control the gas discharge mechanism so as to discharge the gas into the heat insulating sheet TT in a state where the heat insulating sheet TT is folded. Thereby, since the heat insulation sheet swells with gas from the state in which the heat insulation sheet TT is folded, the outside of the spacecraft RV1 can be covered with the heat insulation sheet TT.
  • the probe according to the third embodiment includes an interface connected to a payload or an accessory, and the interface converts a voltage to the payload or the accessory when the interface is connected to the payload or the accessory. And a communication unit for exchanging electrical signals with the payload or accessory when connected to the payload or accessory.
  • This configuration provides power to the payload or accessory and allows the spacecraft to exchange electrical signals with the payload or accessory.
  • FIG. 30 is a perspective view of the spacecraft according to the third embodiment.
  • the probe RV2 includes an antenna AT, a camera CR, a resource search sensor SS, and a drill DR for mining.
  • the payload or accessory according to the present embodiment is, for example, a camera, a resource search sensor, and a mining drill.
  • the resource exploration sensor is a sensor for exploring resources. Resources include minerals and water. Here, as an example, the resource exploration sensor explores water. Note that the payload or accessory is not limited to these and may be any one that is connected to the probe.
  • FIG. 31 is a functional block diagram of the spacecraft according to the third embodiment.
  • the spacecraft RV2 includes a housing HS and three interfaces IF connected to the housing HS.
  • a camera CR, a resource exploration sensor SS, and a mining drill DR are connected to each interface IF.
  • the spacecraft RV2 includes a communication unit CM that communicates with a lander or another spacecraft via an antenna AT, a processor PS that controls the communication unit CM, and a power supply BT.
  • the communication by the communication unit CM may be wired or wireless.
  • FIG. 32 is a functional block diagram of an interface according to the third embodiment.
  • the interface IF includes a communication unit CU, a DC converter unit DCU, and a recording unit MU.
  • the communication unit CU exchanges an electric signal with the payload or the accessory.
  • the payload or accessory is the camera CR
  • a control signal for controlling the camera CR is transmitted to the camera CR via the communication unit CU.
  • the payload or the accessory is the drill DR
  • a control signal for controlling the drill DR is transmitted to the drill DR via the communication unit CU.
  • the DC converter unit DCU converts the voltage of the power supply BT to the payload or the accessory and supplies power.
  • the voltage of the power supply BT is converted into a voltage for the camera CR, and the converted voltage is supplied to the camera CR.
  • the payload or accessory is the resource exploration sensor SS
  • the voltage of the power supply BT is converted into a voltage for the resource exploration sensor SS, and the converted voltage is supplied to the resource exploration sensor SS.
  • the payload or accessory is a drill DR
  • the voltage of the power supply BT is converted into a voltage for the drill DR, and the converted voltage is supplied to the camera CR.
  • the recording unit MU When the recording unit MU is connected to the payload or accessory, the recording unit MU records and / or transfers data acquired by the payload or accessory. For example, the recording unit MU transfers the sensor signal obtained by the resource exploration sensor SS to the processor PS. This sensor signal is transmitted from the communication unit CM to a lander or another probe. In addition, the recording unit MU transfers, for example, a video signal obtained by the camera CR to the processor PS. The video signal represents a still image or a moving image. This video signal is transmitted from the communication unit CM to a lander or another probe.
  • a method of manufacturing a probe component according to the fourth embodiment is a process of manufacturing a probe component by a 3D printer arranged on a celestial body other than the earth (for example, a planet, a satellite, an asteroid, or a comet).
  • the parts of the spacecraft include wheels, housings, antennas, solar panels, instruments, electrical harnesses, and the like.
  • FIG. 33 is a schematic diagram for explaining an example of the manufacturing method according to the fourth embodiment.
  • the 3D printer PR arranged on the moon surface manufactures wheels TY and solar panels SP.
  • the wheel TY and solar panel SP obtained by manufacture are attached to the lander L5.
  • the installation may be performed manually by an astronaut on the moon, by an astronaut operating a robot arm, or by moving a robot arm of the robot. Good.
  • the spacecraft part when a part of the spacecraft breaks down or is damaged, the spacecraft part (for example, the same spacecraft) is melted with the 3D printer using the melted material as a raw material. Such faulty or damaged parts). Thereby, when the parts of the spacecraft are out of order or broken, the spacecraft parts can be manufactured by reusing the spacecraft parts.
  • a 3D printer for example, manufactures the same or failed part of the same spacecraft as a spacecraft part
  • the same part of the spacecraft can be manufactured by reusing the failed or damaged part. it can.
  • you may manufacture the part of another spacecraft for example, a spacecraft or a lander), not only the same part, but another part of a spacecraft.
  • a probe manufacturing method for manufacturing a probe in a celestial body other than the earth uses a step of melting a failed or damaged probe part and a material after melting as a raw material, For example, manufacturing a probe component with a 3D printer on a planet, a satellite, an asteroid, or a comet, and attaching the manufactured probe component to the target probe.
  • the probe can be regenerated by regenerating and replacing the probe component.
  • resources such as planets, satellites, asteroids, or comets
  • non-Earth objects such as natural resources such as minerals and rare metals
  • the spacecraft collects natural resources (for example, minerals) on the moon, and the three-dimensional printer uses the natural resources (for example, minerals) collected on the moon as raw materials.
  • natural resources for example, minerals
  • the three-dimensional printer uses the natural resources (for example, minerals) collected on the moon as raw materials.
  • Manufacture parts when manufacturing a solar panel, a mineral is the silica contained in the regolith of the lunar surface, for example. With this configuration, the parts of the spacecraft can be manufactured at a lower cost. Then, the manufactured probe parts are attached to the target probe. As a result, it is possible to manufacture the probe at a low cost, or to replace a damaged or broken probe component at a low cost.
  • the spacecraft according to the fifth embodiment changes the inclination of the solar panel from the horizontal plane according to the position of the sun. Thereby, the electric power generation amount of a solar panel can be increased.
  • FIG. 34 is a schematic diagram showing an outline of a spacecraft according to the fifth embodiment.
  • the spacecraft RV4 of the fifth embodiment includes a housing HS4, a solar panel SP1 provided on the side surface of the housing HS4, and a drive mechanism DM that changes the inclination of the solar panel SP1.
  • a controller CON that controls the drive mechanism DM.
  • the controller CON changes the inclination with respect to the horizontal plane of the solar panel SP1 according to the time, the position of the sun (for example, the angle of the sun with respect to the horizontal plane), or the amount of power generated by the solar panel SP1.
  • the drive mechanism DM is controlled. Thereby, the inclination based on the horizontal plane of the solar panel SP1 can be changed according to the irradiation angle of sunlight, and the amount of light hitting the solar panel SP1 can be increased. The amount can be increased.
  • the position of solar panel SP1 was provided in the side surface of housing
  • the solar panel SP1 may be wound like a carpet, and the wound solar panel SP1 may be expanded when it is desired to generate power.
  • FIG. 35 is a schematic diagram showing a schematic configuration of the exploration system according to the sixth embodiment.
  • the exploration system S5 includes a lander L11 provided with a reflector RL1, a explorer R11 provided with a reflector RL2, and a solar panel (not shown) on the outer surface of the own aircraft.
  • a spacecraft R12 provided on the surface.
  • the lander L11 and the spacecraft R11 are disposed at positions where the sunlight hits.
  • the spacecraft R12 is searching for a vertical hole in the moon and is located in the shadow area SA, and the spacecraft R12 is in a location where the sun does not hit.
  • the solar light is reflected by the reflector RL1, so that the spacecraft R12 is irradiated with the solar light.
  • the solar light is reflected by the reflector RL2, so that the spacecraft R12 is irradiated with the solar light.
  • electric power is generated by the solar panel of the probe R12, and the probe R12 is driven using the generated electric power.
  • the spacecraft R12 drives a power source (not shown, for example, a motor or an engine) with the generated electric power.
  • a power source not shown, for example, a motor or an engine
  • the spacecraft R11 includes a controller CON.
  • the controller CON determines the direction and / or angle of the reflector RL2 so that the reflected light is irradiated to another probe R12 according to the position of the sun relative to the own aircraft and the position of the other probe R12. May be changed. Thereby, even if the position of the sun and the position of another spacecraft R12 change, the reflected light is irradiated to the spacecraft R12, so that the solar panel of another spacecraft R12 can continue to generate electricity. And the search can be continued using the generated power.
  • the lander L11 includes a controller CON.
  • the controller CON changes the orientation and / or angle of the reflector RL1 so that the reflected light is irradiated to the spacecraft R12 according to the position of the sun relative to the spacecraft and the position of the spacecraft R12. You may do it. Thereby, even if the position of the sun and the position of the spacecraft R12 change, the reflected light is irradiated to the spacecraft R12, so that the solar panel of the spacecraft R12 can continue the power generation, and the generated power The exploration can be continued using.
  • the exploration system S5 includes the reflectors RL1 and RL2 and the explorer R12 having a solar panel.
  • the spacecraft R12 When the spacecraft R12 is in a place where it is not exposed to the sunlight, the sunlight is reflected by the reflector RL1 provided in the lander L11 and / or the reflector RL2 provided in the explorer R11.
  • the spacecraft R12 is irradiated with solar light and is generated by the solar panel of the spacecraft R12. With this configuration, it is possible to generate power even in a place where the sun does not hit the probe R12, and the search can be continued using the generated power.
  • the spacecraft R11 is configured so that the reflector RL2 reflects light and the reflector RL2 is oriented so that sunlight reflected by the reflector RL2 is applied to the solar panel of the object. And / or a controller C1 that controls to change the angle. Even if it is a case where a target object is in a shadow by this structure, since the reflected light can be irradiated to the solar panel of a target object, the solar panel of a target object can generate electric power. In particular, when the object is another probe R12, even if another probe R12 is in the shadow, the reflected light can be applied to the solar panel of the other probe R12. The solar panel of machine R12 can generate electricity and can continue exploration.
  • the target object irradiated with the reflected sunlight was described as the spacecraft R12 as an example, it may be the lander L11.
  • the lander L11 when the lander L11 is located in the shadow area SA, the lander L11 can generate power and continue communication and the like by irradiating the solar panel of the lander L11 with the reflected sunlight. can do.
  • the spacecraft R11 includes a camera that captures another spacecraft R12, and the controller C1 uses the image captured by the camera to convert the sunlight reflected by the reflector RL2 into another spacecraft R12. You may control to change the direction and / or angle of reflector RL2 so that it may follow and irradiate with the solar panel. Thereby, even if another probe R12 moves, reflected light can be irradiated to the solar panel of the probe R12.
  • the lander L11 includes a camera that images the spacecraft R12, and the controller C2 uses the image captured by the camera to convert the sunlight reflected by the reflector RL1 into the sunlight of the spacecraft R12. You may control to change the direction and / or angle of reflecting plate RL1 so that it may irradiate following a panel. Thereby, even if another probe R12 moves, reflected light can be irradiated to the solar panel of the probe R12.
  • the spacecraft emits a camera to a place where it is difficult for the spacecraft to enter (in this case, a vertical hole). Get the captured image.
  • FIG. 36 is a schematic diagram showing a schematic configuration of the exploration system according to the seventh embodiment.
  • the exploration system S8 includes a probe R41 arranged on the moon surface.
  • the probe R41 includes a probe main body B41, a wiring WR, a camera CR connected to the probe main body B41 via the wiring WR, and an injection mechanism IJ for injecting the camera CR.
  • FIG. 36 shows a state in which the camera CR is ejected by the ejection mechanism IJ and landed after the ejection. In this state, the camera CR captures an image and transmits an image obtained by the capture to the probe R41 via the wiring WR.
  • the probe R41 includes the camera CR, the injection mechanism IJ that can inject the camera CR, and the controller CON that controls the camera CR and the injection mechanism IJ.
  • the controller CON controls the camera CR to take a picture at a point where the camera CR has landed after injection, and acquires an image obtained by the photography from the camera.
  • an image at a location where it is difficult for the spacecraft to enter the camera CR is obtained, and thus a location where the spacecraft R41 is difficult to enter (here, a vertical hole) can be observed.
  • the exploration method according to the seventh embodiment includes a step of ejecting the camera CR from the probe R41 having the ejection mechanism IJ, a step of photographing at the point where the camera CR has landed after the ejection, and the camera CR And a step of transmitting the image obtained by the above to the probe R41 via the wiring WR.
  • the spacecraft R41 may further have a winding mechanism for winding the wiring WR.
  • the take-up mechanism may take up the wiring WR so that the camera CR can be stored in the injection mechanism IJ in an injectable manner. Thereby, it can inject
  • the camera CR and the probe main body B41 may each have a wireless communication function, and in that case, an image may be wirelessly transmitted from the camera CR to the probe main body B41.
  • the spacecraft according to the eighth embodiment measures a case, a switching mechanism that switches between an open state that releases heat in the case and a blocking state that blocks heat in the case, and a temperature outside the case. And a processor, and the processor controls the switching mechanism to switch between the open state and the shut-off state according to the temperature measured by the temperature sensor.
  • FIG. 37 is a schematic diagram showing a schematic configuration of the spacecraft according to the eighth embodiment.
  • the probe RV5 includes a housing HS5, a processor PS, a temperature sensor TS, and a switching mechanism SW.
  • the temperature sensor TS measures the temperature outside the housing HS5.
  • the switching mechanism SW switches between an open state in which the heat in the housing HS5 is released and a blocking state in which the heat in the housing HS5 is cut off.
  • the processor PS controls the switching mechanism SW so as to switch between the open state and the shut-off state according to the temperature measured by the temperature sensor TS.
  • FIG. 38A is a schematic perspective view illustrating an example of the switching mechanism SW in the cutoff state.
  • FIG. 38B is a schematic perspective view illustrating an example of the switching mechanism SW in the open state.
  • the switching mechanism SW includes a first frame HM, a second frame IM stacked on the first frame, a shutter frame SF provided on the second frame IM, and a shutter frame.
  • the shutter SH is slidable in the longitudinal direction with respect to the SF.
  • a plurality of rectangular first through holes are formed in the shutter SH at intervals.
  • a plurality of rectangular second through holes are formed at intervals.
  • the second through hole is, for example, approximately the same size as the first through hole.
  • the shutter frame SF When the switching mechanism SW is in the shut-off state, as shown in FIG. 38A, the shutter frame SF has a main body portion (the second through hole is not opened) below the first through hole of the shutter SH. Part) is arranged. Thereby, when the sun hits the spacecraft RV5, heat from the outside is blocked by the main body portion of the shutter frame SF, so that a heat insulating effect can be obtained, and a temperature rise inside the housing HS5 can be suppressed.
  • the shutter frame SF is preferably made of a highly heat-insulating material. Thereby, when the switching mechanism SW is in the cut-off state, the heat insulation effect can be improved. Also, as shown by the arrows in FIG. 38A, the internal gas is blocked by the back surface of the main body portion of the shutter SH (the portion where the first through hole is not opened) and does not escape to the outside.
  • the switching mechanism SW when the switching mechanism SW is in the open state, as shown in FIG. 38B, in the shutter frame SF, the second through hole of the shutter frame SF is disposed below the first through hole of the shutter SH. Thereby, the heat inside the housing HS5 is discharged to the outside of the housing HS5 through the first through hole of the shutter SH and the second through hole of the shutter frame SF.
  • the processor PS may control to switch to the shut-off state as shown in FIG. 38A. Thereby, inflow of the heat
  • the processor PS may control to switch to the open state as shown in FIG. 38B. Thereby, the heat inside the housing HS5 can be discharged as shown by the arrow in FIG. 38B. In this way, the shutter SH is closed when it is hot such as when the sun hits it, and the shutter SH is opened when it is cold such as when the sun does not hit it, so that a change in the temperature inside the housing HS5 can be suppressed. .
  • the temperature sensor TS may measure the temperature inside the housing HS5. Accordingly, the processor PS can control the switching mechanism SW so as to switch between the open state and the shut-off state according to the temperature inside the housing HS5.
  • the processor PS may control the switching mechanism SW so as to switch between an open state and a shut-off state according to a preset month cycle.
  • the processor PS may control the switching mechanism SW so as to switch to the cut-off state during a period when sunlight falls on the moon, and may control the switching mechanism SW so as to switch to an open state when the moon does not receive sunlight. Good.
  • the spacecraft according to the ninth embodiment folds the legs that support the wheels. With this configuration, the center of gravity is lowered, so that when the vibration during transportation or the like is large, the vibration of the spacecraft due to the vibration can be reduced and the damage due to the vibration can be suppressed. Further, when the spacecraft or the lander is loaded with the spacecraft as a payload, the height can be reduced by folding the legs, so that the space occupied by the spacecraft when loaded is reduced.
  • FIG. 39A is a schematic side view of the spacecraft when the legs are not folded.
  • FIG. 39B is a schematic side view of the spacecraft when the legs are folded.
  • FIG. 39C is a schematic perspective view of a leg portion and a support column. As shown in FIGS. 39A to 39C, the housing HS, the support column PL connected to the housing HS, the two pairs of legs LG1 and LG2 whose one ends are connected to the support PL, and the legs LG1 and LG2, respectively. And wheels TY1, TY2 connected to the other end.
  • the legs LG1 and LG2 are supported by the holding structure HD.
  • the two pairs of legs LG1 and LG2 are configured to be rotatable with respect to the support column PL so that the angle between the two pairs of legs LG1 and LG2 is variable.
  • the two pairs of legs LG1, LG2 are rotated to the vicinity of the lower surface of the housing HS to open the angle between the two pairs of legs LG1, LG2.
  • LG1 and LG2 can be folded.
  • the legs LG1 and LG2 may be made of, for example, flexible carbon fiber reinforced plastic (Carbon Fiber Reinforced Plastic: CFRP).
  • the spacecraft according to the ninth embodiment is connected to the housing, the support column connected to the housing, the two pairs of legs connected at one end to the support column, and the other ends of the legs.
  • the two pairs of legs are configured to be rotatable with respect to the support so that the angle between the two pairs of legs is variable.
  • the casing of the spacecraft according to each embodiment may include graphene or graphene fiber as a material. Part of the material of the housing may be used, or all of the material may be used. Thereby, the heat insulation of a housing
  • casing can be improved.
  • the present disclosure is not limited to the above-described embodiment as it is, and can be embodied by modifying the constituent elements without departing from the scope in the implementation stage.
  • various inventions can be formed by appropriately combining a plurality of components disclosed in the embodiment. For example, some components may be deleted from all the components shown in the embodiment.
  • constituent elements over different embodiments may be appropriately combined.
  • AD1, AD2, AD3, AD4 A / D converter AT antenna AT1 first antenna AT2 second antenna B1, B2 bolt B41 probe body BAT battery BB battery board BC rear camera BHP bearing hold plate BP rear plate BR1, BR2 bearing BS Bearing spacer BT Power source C1, C2, CON Controller CB Camera board CC1 First communication controller CC2 Second communication controller CDC Charge / discharge circuit CM communication unit CMC1 First camera controller CMC2 Second camera controller CO Notch CR Camera CU Communication unit D1 Diode DCU DC converter unit DM Drive mechanism DP Bottom plate DR Drill DS Distance sensor E Ground station ET Earth FC Front camera FP Front plate FW1, F W2, RW1, RW2 Wheel GB Gearbox H1 Through-hole HB Hub HD Holding structure HC Clamp HE1 Hole HI Thermal insulation HL First hole HL2 Second hole HM First frame HS, HS2, HS4 Housing IF interface IM First 2 frames L Lander L5, L11 Lander LC Left side camera LG1, LG2 Leg LP

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Abstract

L'invention concerne une sonde pouvant se déplacer, ladite sonde comportant : des roues ; une première caméra disposée en regard de la direction dans laquelle la sonde peut se déplacer ; et une seconde caméra disposée en regard des directions ne comprenant pas la direction dans laquelle la sonde peut se déplacer. La lentille de la première caméra et/ou de la seconde caméra est dirigée vers le bas à partir de la ligne horizontale, et les roues sont comprises dans le champ visuel de la première caméra et/ou le champ visuel de la seconde caméra.
PCT/JP2017/028682 2016-08-10 2017-08-08 Sonde, procédé de fabrication d'élément de sonde et procédé de fabrication de sonde Ceased WO2018030368A1 (fr)

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CN109484673A (zh) * 2018-12-24 2019-03-19 深圳航天东方红海特卫星有限公司 一种载荷平台分离式遥感微小卫星构型及其装配方法

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CN110626523A (zh) * 2019-10-18 2019-12-31 中国矿业大学(北京) 月球区域观测与通信系统及弹跳装置
CN112249367B (zh) * 2020-10-13 2022-04-05 哈尔滨工业大学 一种小行星探测机动巡视装置

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CN108725845A (zh) * 2018-08-16 2018-11-02 重庆大学 着陆缓冲与隔振一体化悬架
CN108725845B (zh) * 2018-08-16 2020-11-24 重庆大学 着陆缓冲与隔振一体化悬架
CN109484673A (zh) * 2018-12-24 2019-03-19 深圳航天东方红海特卫星有限公司 一种载荷平台分离式遥感微小卫星构型及其装配方法

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