Spacecraft and method for manufacturing same

Abstract

[Problem] To provide a spacecraft capable of properly controlling the attitude thereof, and a method for manufacturing same. [Solution] A spacecraft 2 comprises: a spacecraft body 10; and a movement control device 12 installed on the spacecraft body 10 and having a deployment membrane 14 that is deployed to act with atmospheric resistance, wherein the movement control device 12 is installed on the spacecraft body 10 such that the geometric center of gravity of the deployment membrane 14 in a deployed state is offset from the center of gravity of the spacecraft body 10 in the direction of a roll axis of the spacecraft. The moment of inertia of the spacecraft 2 around the roll axis is greater than the moment of inertia around a pitch axis and the moment of inertia around a yaw axis.

Description

Spacecraft and its manufacturing method

[0001] The present invention relates to a spacecraft equipped with a movement control device and a method for manufacturing the same.

[0002] In the space industry, spacecraft such as artificial satellites, rockets, and space stations are operated in Earth orbit. While orbiting the Earth at a predetermined altitude, spacecraft collect and monitor information on the Earth and outer space, conduct experiments in the space environment, communicate with the Earth, and transport goods.

[0003] If a spacecraft continues to orbit the Earth after completing its designated mission, it may become space debris and interfere with the operation of other spacecraft. Therefore, it is necessary to de-orbit the spacecraft after its operation has ended. However, attempting to de-orbit a spacecraft using a propulsion mechanism such as an engine requires extensive movement control of the spacecraft, which is time-consuming and costly. Furthermore, a spacecraft may not have enough energy remaining after its operation has ended to de-orbit the spacecraft.

[0004] Therefore, spacecraft are sometimes equipped with a movement control device (post-orbital disposal device) for de-orbiting the spacecraft. Such a movement control device is also called a PMD (Post Mission Disposal) device, and operates independently of the spacecraft itself to de-orbit the spacecraft.

[0005] For example, a movement control device is used that stretches a tape-like conductive tether into space to decelerate a spacecraft (see Patent Document 1). When the conductive tether extended from the movement control device crosses the Earth's magnetic field, an induced electromotive force is generated in the conductive tether, causing an induced current to flow. The mutual induction between the induced current and the Earth's magnetic field generates a Lorentz force in the conductive tether. This Lorentz force acts in the opposite direction to the spacecraft's direction of travel, causing the spacecraft to decelerate and leave its orbit.

[0006] However, it is difficult to control the movement of a spacecraft using a motion control device to correctly extend a long conductive tether in a predetermined direction. Furthermore, to precisely control the movement of a spacecraft using a conductive tether, the electron emitter at the end of the conductive tether must operate stably at all times. This requires high reliability in the control of the motion control device, which also consumes a lot of power.

[0007] Therefore, a movement control device that uses atmospheric resistance to decelerate a spacecraft has also been proposed (see Patent Document 2). This type of movement control device has a deployment membrane that is acted upon by atmospheric resistance, and the deployment membrane is deployed and spread when the spacecraft de-orbits. As a result, atmospheric resistance acts on the deployment membrane, decelerating the spacecraft and allowing it to depart from orbit.

[0008] JP 2021-115713 A JP 2022-143395 A

[0009] By installing a movement control device on a spacecraft, it becomes possible to decelerate the spacecraft and depart from Earth's orbit, as described above, or to temporarily change the direction or speed of the spacecraft. Furthermore, by adopting a method of controlling the movement of a spacecraft using atmospheric resistance acting on the deployment membrane, the movement control of the spacecraft is simplified compared to when a conductive tether is used.

[0010] To control the movement of a spacecraft, it is preferable for the spacecraft to maintain a constant attitude as much as possible. However, a spacecraft in an orbit does not always move forward with a constant attitude, and some external force may cause it to tilt or rotate. As a result, the movement control capability of the movement control device is reduced, making it difficult to smoothly perform operations such as deorbiting the spacecraft.

[0011] Furthermore, when the movement control device deploys the deployment membrane, the mass properties of the entire spacecraft fluctuate. This makes it difficult for the spacecraft to maintain a constant attitude, and it may enter a state in which it moves while rotating (tumbling mode). When the spacecraft enters tumbling mode, it becomes more difficult for the movement control device to control the movement of the spacecraft. This may make it impossible to execute deorbit of the spacecraft itself, or even if deorbit is possible, it may become impossible to predict or control the timing of the spacecraft's arrival on the ground or the point of impact. Furthermore, when a flying object such as space debris is approaching the spacecraft, it becomes difficult to control the spacecraft to temporarily change its direction of travel or speed to avoid a collision with the flying object.

[0012] The present invention has been made in view of the above problems, and has as its object to provide a spacecraft capable of appropriately controlling the attitude of the spacecraft, and a method for manufacturing the same.

[0013] According to one aspect of the present invention, there is provided a spacecraft comprising: a spacecraft body; and a movement control device attached to the spacecraft body, the movement control device having a deployment membrane that is deployed so that atmospheric resistance acts on it, wherein the movement control device is attached to the spacecraft body so that the geometric center of gravity of the deployment membrane in a deployed state is offset in the direction of the roll axis of the spacecraft from the center of gravity of the spacecraft body, and wherein the moment of inertia about the roll axis of the spacecraft is greater than the moment of inertia about the pitch axis and the moment of inertia about the yaw axis.

[0014] According to another aspect of the present invention, there is provided a spacecraft comprising: a spacecraft body; and a movement control device attached to the spacecraft body, the movement control device having a deployment membrane that is deployed so that atmospheric resistance acts on it, wherein the movement control device is attached to the spacecraft body so that the geometric center of gravity of the deployment membrane in a deployed state is offset from the center of gravity of the spacecraft body in the direction of the spacecraft's roll axis, and wherein the moment of inertia of the spacecraft about the roll axis is greater than the moment of inertia about the yaw axis.

[0015] According to yet another aspect of the present invention, there is provided a spacecraft comprising: a spacecraft body; and a movement control device attached to the spacecraft body, the movement control device having a deployable membrane that is deployed so that atmospheric resistance acts on it, wherein the movement control device is attached to the spacecraft body so that the geometric center of gravity of the deployable membrane in a deployed state is offset in the direction of the roll axis of the spacecraft from the center of gravity of the spacecraft body, and wherein the moment of inertia of the spacecraft about the roll axis is greater than the moment of inertia about the yaw axis, and the moment of inertia about the pitch axis is greater than the moment of inertia about the roll axis.

[0016] Preferably, the unfolded membrane in the unfolded state has an inclined surface that forms an acute angle with the roll axis.

[0017] Preferably, the movement control device further includes an angle adjustment mechanism capable of adjusting the angle between the deployment membrane and the roll axis of the spacecraft in a deployed state.

[0018] Preferably, the spacecraft body or the movement control device has an electronic computer that controls the operation of the angle adjustment mechanism.

[0019] Preferably, the movement control device further comprises one or both of a solar cell and a battery.

[0020] Preferably, the movement control device causes the spacecraft body to leave the orbit by deploying the deployment membrane.

[0021] According to yet another aspect of the present invention, there is provided a method for manufacturing a spacecraft, comprising: preparing a spacecraft main body; and a movement control device having a deployable membrane that is deployed so that atmospheric resistance acts on it; and attaching the movement control device to the spacecraft main body so that the geometric center of gravity of the deployable membrane in a deployed state is offset from the center of gravity of the spacecraft main body in the roll axis direction.

[0022] Preferably, the movement control device is attached to the spacecraft body so that the deployment membrane in the deployed state forms an acute angle with respect to the rear of the roll axis.

[0023] According to yet another aspect of the present invention, there is provided a spacecraft comprising: a spacecraft main body; and a movement control device attached to the spacecraft main body, the movement control device having a deployment membrane that is deployed so that atmospheric resistance acts on it, wherein the movement control device is attached to the spacecraft main body so that the deployment membrane in a deployed state has an inclined surface that forms an acute angle with respect to the roll axis, and wherein the moment of inertia of the spacecraft about the roll axis is greater than the moment of inertia about the pitch axis and the moment of inertia about the yaw axis.

[0024] According to yet another aspect of the present invention, there is provided a spacecraft comprising: a spacecraft body; and a movement control device attached to the spacecraft body, the movement control device having a deployment membrane that is deployed so that atmospheric resistance acts on the deployment membrane, wherein the movement control device is attached to the spacecraft body so that the deployment membrane in a deployed state has an inclined surface that forms an acute angle with respect to the roll axis, and wherein the moment of inertia of the spacecraft about the roll axis is greater than the moment of inertia about the yaw axis.

[0025] According to yet another aspect of the present invention, there is provided a spacecraft comprising: a spacecraft body; and a movement control device attached to the spacecraft body, the movement control device having a deployment membrane that is deployed so that atmospheric resistance acts on it, wherein the deployment membrane in a deployed state has an inclined surface that forms an acute angle with respect to the roll axis, and wherein the moment of inertia of the spacecraft about the roll axis is greater than the moment of inertia about the yaw axis, and the moment of inertia about the pitch axis is greater than the moment of inertia about the roll axis.

[0026] According to yet another aspect of the present invention, there is provided a method for manufacturing a spacecraft, comprising: preparing a spacecraft main body; and a movement control device having a deployable membrane that is deployed so that atmospheric resistance acts on the spacecraft main body; and attaching the movement control device to the spacecraft main body so that the deployable membrane in a deployed state has an inclined surface that forms an acute angle with respect to a roll axis.

[0027] The spacecraft and its manufacturing method of the present invention provide a spacecraft in which the geometric center of gravity of the deployment membrane is offset from the center of gravity of the spacecraft with respect to the roll axis, and a spacecraft in which the deployment membrane in the deployed state has an inclined surface that forms an acute angle with the roll axis. As a result, when the attitude of the spacecraft changes, an attitude restoring torque is applied to correct the attitude change, making it possible to appropriately control the attitude and movement of the spacecraft.

[0028] FIG. 1 is a perspective view showing a spacecraft according to a first embodiment. FIG. 2(A) is a front view showing the movement control device with the deployment membrane retracted, and FIG. 2(B) is a front view showing the movement control device with the deployment membrane deployed. FIG. 3(A) is a diagram schematically showing an example of a state in which atmospheric resistance is applied to the deployment membrane when the spacecraft according to the first embodiment is flying with its roll axis parallel to the direction of travel. FIG. 3(B) is a diagram schematically showing an example of a state in which atmospheric resistance is applied to the deployment membrane when the spacecraft according to the first embodiment is flying with its roll axis tilted relative to the direction of travel. FIG. 4(A) is a diagram schematically showing an example of a state in which atmospheric resistance is applied to the deployment membrane when the spacecraft according to the reference example is flying with its roll axis parallel to the direction of travel. FIG. 4(B) is a diagram schematically showing an example of a state in which atmospheric resistance is applied to the deployment membrane when the spacecraft according to the reference example is flying with its roll axis tilted relative to the direction of travel. FIG. 5 is a diagram schematically illustrating another example of a state in which atmospheric resistance is applied to the deployable membrane when the spacecraft of the first embodiment is flying with its roll axis tilted relative to the direction of travel. FIG. 6 is a diagram schematically illustrating another example of a state in which atmospheric resistance is applied to the deployable membrane when the spacecraft of the reference example is flying with its roll axis tilted relative to the direction of travel. FIG. 7 is a perspective view of a spacecraft according to a second embodiment. FIG. 8 is a diagram schematically illustrating an example of a state in which atmospheric resistance is applied to the deployable membrane when the spacecraft of the second embodiment is flying with its roll axis tilted relative to the direction of travel. FIG. 9 is a perspective view showing the configuration of a movement control device provided in the spacecraft of the second embodiment. FIG. 10 is a diagram schematically illustrating another example of a state in which atmospheric resistance is applied to the deployable membrane when the spacecraft of the second embodiment is flying with its roll axis tilted relative to the direction of travel. FIG. 11 is a perspective view of a spacecraft according to a third embodiment. FIG. 12 is a perspective view of a spacecraft according to a fourth embodiment. Fig. 13 is a side view of the spacecraft of Fig. 12. Fig. 14 is a perspective view showing a spacecraft according to a fifth embodiment. Fig. 15 is a side view of the spacecraft of Fig. 14.

[0029] First Embodiment An embodiment according to one aspect of the present invention will be described below with reference to the accompanying drawings. First, a configuration example of a spacecraft according to this embodiment will be described. Fig. 1 is a perspective view showing a spacecraft 2.

[0030] The spacecraft 2 is a flying vehicle that orbits a celestial body 4 such as the Earth along a circular orbit 6. For example, the spacecraft 2 may be an artificial satellite, a rocket, a transport vehicle, a space probe, a space station, or the like, and may be equipped with a propulsion mechanism such as a chemical propulsion engine or an electric propulsion engine. However, there is no limitation on the type of the spacecraft 2, and the spacecraft 2 may also be a flying vehicle that does not have a propulsion mechanism.

[0031] The orbit 6 is a travel path of the spacecraft 2 set outside the celestial body 4. For example, the orbit 6 is a substantially circular or elliptical path set so that the distance (altitude) from the surface (earth's surface) of the celestial body 4 is approximately constant. The altitude of the orbit 6 is set depending on the type and operational purpose of the spacecraft 2, and is, for example, between 200 km and 1000 km.

[0032] Figure 1 shows three mutually perpendicular axes: the X-axis, the Y-axis, and the Z-axis. The X-axis corresponds to the tangential direction of the orbit 6 and indicates the direction of travel of the spacecraft 2. The Z-axis corresponds to the normal direction of the orbit 6 and indicates the altitude direction of the spacecraft 2. The positive direction of the Z-axis (upward in Figure 1) corresponds to the zenith direction, which is the direction away from the celestial body 4. On the other hand, the negative direction of the Z-axis (downward in Figure 1) corresponds to the nadir direction, which is the direction toward the center (geocenter) of the celestial body 4.

[0033] The spacecraft 2 also has a roll axis, a pitch axis, and a yaw axis, which are three mutually perpendicular axes that pass through the center of gravity of the spacecraft 2. The spacecraft 2 is designed assuming that the roll axis, pitch axis, and yaw axis are arranged parallel to the X axis, Y axis, and Z axis, respectively. Therefore, the roll axis direction of the spacecraft 2 corresponds to the direction assumed to be the direction of travel of the spacecraft 2. When the spacecraft 2 is moved in the circular orbit 6, the attitude of the spacecraft 2 is controlled so that the roll axis, pitch axis, and yaw axis are parallel to the X axis, Y axis, and Z axis, respectively.

[0034] For ease of explanation, it is assumed below that the roll axis, pitch axis, and yaw axis of the spacecraft 2 placed on the circular orbit 6 are parallel to the X axis, Y axis, and Z axis, respectively. However, when the spacecraft 2 actually moves on the circular orbit 6, the attitude of the spacecraft 2 fluctuates, and the roll axis, pitch axis, and yaw axis may not be parallel to the X axis, Y axis, and Z axis, respectively.

[0035] During operation, the spacecraft 2 orbits the celestial body 4 along the circular orbit 6 and performs a pre-specified mission (collection and monitoring of information on the Earth and outer space, experiments in the space environment, communication with the Earth, transport of goods, etc.). When the mission is completed and the operation of the spacecraft 2 ends, the spacecraft 2 leaves the circular orbit 6 and enters a departure orbit 8, gradually approaching the celestial body 4 while orbiting it. The spacecraft 2 then enters the atmosphere and burns up, or reaches the celestial body 4 and is recovered. The period required for the spacecraft 2 to leave its orbit varies depending on the type, altitude, etc. of the spacecraft 2, and could be several hours, several weeks, several months, several years, or even more than 10 years.

[0036] Next, the detailed configuration of the spacecraft 2 will be described. The spacecraft 2 comprises a spacecraft main body (hereinafter simply referred to as "main body") 10 on which the main functions of the spacecraft 2 are mounted, and a movement control device 12 attached to the main body to control the movement of the main body 10. The movement control device 12 controls the direction of travel and movement speed of the spacecraft 2 by utilizing the resistance of the atmosphere around the celestial body 4. In particular, when the spacecraft 2 is to leave the orbit 6, the movement control device 12 functions as an orbit departure device.

[0037] The movement control device 12 is equipped with multiple deployment membranes (resistive membranes) 14 that are deployed so that atmospheric resistance acts on them. The multiple deployment membranes 14 are arranged at approximately equal angular intervals around the roll axis and are disposed symmetrically with respect to the roll axis of the spacecraft 2. For example, if the movement control device 12 is equipped with two sets of deployment membranes 14 as shown in FIG. 1 , the two sets of deployment membranes 14 are attached to both ends of the main body 10 in the yaw axis direction, i.e., the upper end and the lower end. As a result, the two sets of deployment membranes 14 are disposed symmetrically at 180° intervals around the roll axis.

[0038] However, there is no limit to the number of deployable membranes 14. For example, if the movement control device 12 is equipped with four sets of deployable membranes 14, the four sets of deployable membranes 14 are arranged at 90° intervals around the roll axis, and each deployable membrane 14 is arranged symmetrically with respect to the other deployable membranes 14 with the roll axis as the center. Alternatively, one deployable membrane 14 may be provided so as to surround the main body 10 around the roll axis, or three deployable membranes 14 may be arranged so that their axes form 120° with each other when viewed from the roll axis direction.

[0039] Furthermore, the arrangement of the deployment membranes 14 does not necessarily need to be symmetrical with respect to the roll axis, nor does it need to be arranged at equal intervals around the roll axis. It is sufficient that the position of the geometric center of gravity of the deployed deployment membrane 14 coincides with the position of the center of gravity of the entire spacecraft 2 when viewed from the roll axis direction.

[0040] In the following description, the geometric center of gravity of the deployable membrane 14 refers to the "position of the center of gravity of all deployable membranes 14 provided on the spacecraft 2." For example, if the spacecraft 2 has two deployable membranes 14 symmetrically arranged around the main body 10, the geometric center of gravity of the deployable membranes 14 is located on an imaginary line segment connecting the centers of gravity of the deployable membranes 14.

[0041] The deployable membrane 14 comprises a film (sheet) capable of receiving atmospheric molecules and a pair of fixing members 18 that fix the film in a taut state. The film constitutes the membrane surface 16 that receives atmospheric resistance. The fixing members 18 correspond to booms that support the membrane surface 16 formed by the film. Both side ends of the membrane surface 16 are fixed to the fixing members 18.

[0042] There are no restrictions on the material, shape, dimensions, etc. of the film that constitutes the film surface 16, as long as the film surface 16 can withstand atmospheric resistance. For example, the film can be formed by covering the surface of a resin film with a conductive thin film. This prevents the resin film from being exposed to atomic oxygen (AO) present in the orbit 6 and from deteriorating. Furthermore, by forming a conductive thin film on the resin film to impart conductivity to the film, damage to the film due to charging (charge-up) and discharge can be avoided.

[0043] The film includes a resin film (thickness: about 12.5 μm) made of polyimide or the like, and a conductive film made of aluminum, ITO (indium tin oxide), or the like that covers the surface of the resin film, and is formed in a triangular, trapezoidal, or fan shape. The area of ​​the film surface 16 is, for example, 100 m ２ However, the material, shape, size, etc. of the film that constitutes the membrane surface 16 can be selected appropriately depending on the weight, shape, size, altitude, etc. of the spacecraft 2.

[0044] The fixing member 18 is a member that can maintain a columnar shape and may be capable of deformation or expansion and contraction. For example, a cylindrical member made of carbon fiber reinforced plastics (CFRP) can be used as the fixing member 18. The thickness of the fixing member 18 (the difference between the outer diameter and the inner diameter) can be set to about 0.12 mm, the diameter of the fixing member 18 to about 30 mm, and the length of the fixing member 18 to about 15.2 m. However, the material, shape, size, etc. of the fixing member 18 can be selected appropriately depending on the shape, size, etc. of the membrane surface 16.

[0045] Alternatively, the fixing member 18 may be a cylindrical member (inflatable tube) that expands when filled with gas and is fixed into a columnar shape. The inflatable tube is a deformable hollow cylindrical member, and is formed by rolling a film such as a resin film, a metal film, or a carbon film into a cylindrical shape.

[0046] When the fixing member 18 is an inflatable tube, the movement control device 12 is equipped with a gas supply source (not shown) that supplies gas (air, nitrogen gas, etc.) to the fixing member 18, and an electromagnetic valve (not shown) that controls the supply of gas from the gas supply source to the fixing member 18. By supplying gas from the gas supply source to the inside of the fixing member 18 via the electromagnetic valve, the fixing member 18 expands, and the shape of the fixing member 18 is fixed to a columnar shape.

[0047] During operation of the spacecraft 2, the membrane 16 and the fixing members 18 are housed in the movement control device 12, for example, in a folded state or in a state wound around a core material. When operation of the spacecraft 2 ends, the pair of fixing members 18 are launched from the movement control device 12 into space, and the membrane 16 fixed to the pair of fixing members 18 spreads out and becomes taut. As a result, the atmosphere around the celestial body 4 collides with the membrane 16, and atmospheric resistance acts on the membrane 16 in the direction opposite to the traveling direction of the spacecraft 2. As a result, the spacecraft 2 decelerates and leaves the circular orbit 6, entering the escape orbit 8.

[0048] The fixing member 18 may be an insulator or a conductor. If the fixing member 18 is conductive, a current can be passed through the fixing member 18 by induced electromotive force or an external power source. This makes it possible to use the Lorentz force generated by the mutual induction between the current flowing through the fixing member 18 and the magnetic field of the celestial body 4 to control the movement of the spacecraft 2.

[0049] There is no limitation on the method for imparting conductivity to the fixing member 18. For example, a conductive thin film may be provided on the surface or inside of the fixing member 18, or a conductive material may be contained in the fixing member 18. As the conductive thin film, for example, a thin film containing a metal such as aluminum, nickel, or chromium may be used. Furthermore, as the conductive material, for example, a fiber containing carbon fiber may be used.

[0050] 2(A) is a front view showing the movement control device 12 with the deployable membrane 14 stored therein. The movement control device 12 includes a storage unit 20 that stores the deployable membrane 14. For example, the storage unit 20 is a box-shaped container that stores the membrane surface 16 and the fixing member 18 in a folded or rolled state. There are no limitations on the shape or size of the storage unit 20 as long as the deployable membrane 14 can be stored therein.

[0051] 2(B) is a front view showing the movement control device 12 with the deployment membrane 14 deployed. When the movement control device 12 is activated, the pair of fixing members 18 are ejected from the housing 20 so as to extend in a predetermined direction due to the action of their own elastic energy, etc. For example, the pair of fixing members 18 extend in a direction inclined with respect to the yaw axis (Z axis) so that their tips are spaced apart along the pitch axis (Y axis). When the shape of the fixing members 18 is fixed in a columnar shape, the membrane surface 16 attached to the fixing members 18 becomes taut, and the deployment membrane 14 is deployed.

[0052] The timing for releasing the spacecraft 2 from the orbit 6 is set appropriately depending on the content of the mission to be executed by the spacecraft 2 and the lifespan of the spacecraft 2. For example, the deployable membrane 14 is deployed when a command to deploy the deployable membrane 14 is input to the mobile control device 12 from the main body 10 of the spacecraft 2 or from the ground. The mobile control device 12 may also deploy the deployable membrane 14 when the operation time or flight time of the spacecraft 2 reaches a predetermined time limit. Furthermore, the mobile control device 12 may deploy the deployable membrane 14 when a signal is received notifying that the operation of the main body 10 of the spacecraft 2 has stopped.

[0053] Here, in particular, in the spacecraft 2 of the first embodiment, as shown in FIG. 1, the geometric center of gravity of the entire deployable membrane 14 deployed around the roll axis (G Ｗ The position of the center of gravity of the spacecraft 2 (shown as G in the figure) is Ｓ The spacecraft 2 and the deployment membrane 14 are designed so that they are offset in the roll axis direction from the roll axis (indicated by the symbol ). By doing so, the attitude of the spacecraft 2 during flight is maintained such that the roll axis is aligned with the X-axis direction. The effects of this attitude maintenance will be described below.

[0054] For the sake of convenience, we will refer to the "center of gravity G of spacecraft 2" here. Ｓ " refers to the "center of gravity of the entire spacecraft 2 including the deployment membrane 14 when the deployment membrane 14 is deployed." Ｓas "the center of gravity of the spacecraft main body 10 excluding the deployment membrane 14," the effects described below are basically the same. This is because, whichever definition is adopted, the geometric center of gravity of the deployment membrane 14 will be offset in the roll axis direction from the center of gravity of the entire spacecraft 2, including the deployment membrane 14.

[0055] Fig. 3(A) is a diagram schematically illustrating an example of a state in which atmospheric resistance is applied to the deployable membrane when the spacecraft 2 of the first embodiment is flying with its travel direction and roll axis parallel. Fig. 4(A) is a diagram similarly illustrating an example of a state in which atmospheric resistance is applied to the deployable membrane when the spacecraft is flying with its travel direction and roll axis parallel, but shows a spacecraft 2' as a reference example for comparison with the spacecraft 2 in Fig. 3(A) in which the geometric center of gravity of the deployable membrane 14 is not offset from the center of gravity of the entire spacecraft (the positions of the centers of gravity of the deployable membrane and the entire spacecraft are aligned).

[0056] In Figures 3(A) and 4(A), two deployment membranes 14 are deployed symmetrically around the roll axis of the spacecraft 2 (2') along the yaw axis (in directions forming 180° with respect to each other when viewed from the roll axis direction), and this state is viewed from a direction (pitch axis (Y axis) direction) perpendicular to the roll axis and yaw axis (lines connecting the centers of gravity of the two deployment membranes 14).

[0057] In the spacecraft 2 shown in FIG. 3A, the center of gravity G of the spacecraft 2 is the same as that of the spacecraft 2 shown in FIG. Ｓ and the geometric center of gravity G of the two deployed membranes 14. Ｗ and are offset with respect to the roll axis, and the geometric center of gravity G Ｗ is the center of gravity G Ｓ 4A, the center of gravity G of the spacecraft 2' is located at the rear of the spacecraft 2' in the direction of travel along the roll axis (X axis). Ｓ and the geometric center of gravity G of the two deployed membranes 14. Ｗ and are not offset but coincide with each other.

[0058] During flight of the spacecraft 2, 2', the deployable membranes 14 are deployed to form a plane that intersects with the direction of travel (X-axis direction) of the spacecraft 2, 2', so each deployable membrane 14 receives a resistance force from the atmosphere in a direction from front to rear with respect to the direction of travel. When the two deployable membranes 14 are deployed symmetrically with respect to the roll axis along the yaw axis, each deployable membrane 14 located on the left and right in the figure receives a substantially equal resistance force F from the atmosphere. Ｌ , F Ｒ Receive.

[0059] In the spacecraft 2' of the reference example shown in FIG. 4A, when the two deployment membranes 14 are deployed symmetrically with respect to the roll axis along the yaw axis and the roll axis of the spacecraft 2' is parallel to the X axis, the center of gravity G Ｓ Considering the torque generated around the resistance force F Ｌ , F Ｒ The torques due to the center of gravity G act in opposite directions. Ｗ from the center of gravity G of each deployment membrane 14 Ｌ , G Ｒ If the distance to each point is D, then the center of gravity G Ｓ Around Ｌ , D x F Ｒ Since the torques act in opposite directions, these torques cancel each other out and the spacecraft 2' does not rotate.

[0060] The same is true for the spacecraft 2 of the first embodiment shown in FIG. Ｓ and the geometric center of gravity G of the deployed membrane 14 Ｗ and are offset with respect to the roll axis, the resistance forces F that the left and right deployable membranes 14 receive from the atmosphere along the X-axis direction are respectively Ｌ , F Ｒ Then, the center of gravity G of Spacecraft 2 Ｓ and the center of gravity G of each deployment membrane 14 Ｌ , G Ｒ Component F perpendicular to the line segment connecting Ｌｔ , F Ｒｔ acts as a torque that tries to rotate the spacecraft 2 around the pitch axis (Y axis). Ｓ From the center of gravity G Ｌ , G Ｒ If the distance to each point is D, then the center of gravity GＳ Around Ｌｔ , D x F Ｒｔ Since the torques of the two axes act in opposite directions, they cancel each other out, and the spacecraft 2 does not rotate.

[0061] Next, consider the case where the roll axis of the spacecraft 2, 2' is tilted with respect to the direction of travel (X-axis direction). Ｓ , G Ｗ This diagram shows an example of a state in which atmospheric resistance is applied to the deployment membrane when a reference spacecraft 2', which is not offset from the other spacecraft, is flying with its roll axis tilted relative to the direction of travel.

[0062] In Figure 4B, the spacecraft 2' is tilted around the pitch axis (Y axis) by a small angle φ with respect to the X-axis direction and the Z-axis direction. In this case, the resistance force F that one of the deployment membranes 14 receives from the atmosphere is Ｌ Torque D x F Ｌｔ and the resistance force F that the other deployment film 14 receives from the atmosphere. Ｒ Torque D x F Ｒｔ Therefore, if an external force acts on the spacecraft 2' from the state shown in Figure 4(A), causing the spacecraft 2' to rotate and tilt by an angle φ as shown in Figure 4(B), the rotation of the spacecraft 2' will be affected by the resistance force F that the deployment membrane 14 receives from the atmosphere. Ｌ , F Ｒ It will not be stopped by

[0063] Here, the center of gravity G Ｓ , G Ｗ Consider a case where the spacecraft 2 of the first embodiment, in which the roll axes are offset from each other, is tilted with respect to the X axis in the same manner. Figure 3(B) is a diagram schematically showing an example of a state in which resistance from the atmosphere is applied to the deployable membranes when the spacecraft 2 of the first embodiment is flying with the roll axis tilted with respect to the direction of travel. In the spacecraft 2 of the first embodiment, as already mentioned, the center of gravity G of each deployable membrane 14 Ｌ , G Ｒ is the center of gravity G of Spacecraft 2 Ｓ It is located further back.

[0064] As shown in FIG. 3B, if the spacecraft 2 is tilted around the pitch axis (Y axis) by an angle φ with respect to the X-axis direction and the Z-axis direction, the resistance force F acting on the deployment membrane 14 from the atmosphere will be Ｌ , F Ｒ The center of gravity G of Spacecraft 2 Ｓ The component generated around the resistance force F Ｌ , F Ｒ Of these, the center of gravity G of Spacecraft 2 Ｓ and the center of gravity G of each deployment membrane 14 Ｌ , G Ｒ (component perpendicular to the line segment connecting Ｌｔ , F Ｒｔ are not equal. If the resistance force F Ｌ , F Ｒ are equal to each other, these resistance forces F Ｌ , F Ｒ The position where the force is applied is not symmetrical with respect to the X axis, so the spacecraft 2 is Ｓ The component F of the direction of rotation around Ｌｔ , F Ｒｔ is different.

[0065] The two deployment membranes 14 are aligned at the center of gravity G Ｌ , G Ｒ The position of the center of gravity G of Spacecraft 2 Ｓ As a result, as shown in FIG. 3A, the center of gravity G of the spacecraft 2 is Ｓ The center of gravity G of each deployable membrane 14 as seen from Ｌ , G Ｒ The positions of the center of gravity G are respectively located rearward by an amount corresponding to the angle α with respect to the yaw axis. Ｓ and center of gravity G Ｌ , G Ｒ If the distance between each is D, then the center of gravity G Ｌ , G Ｒ is the center of gravity G Ｓ (The roll axis is positioned behind the roll axis by an angle D sin α.) If the spacecraft 2 is tilted from this state by a small angle φ around the pitch axis (Y axis) as shown in FIG. 3B, the center of gravity G of one of the deployment membranes 14 will be Ｌ The position of the center of gravity G of Spacecraft 2 ＳThe center of gravity G of the other deployable membrane 14 is located rearward by D sin (α + φ) with respect to the X axis when viewed from the Ｒ The position of the center of gravity G of Spacecraft 2 Ｓ When viewed from the front, it will be positioned a distance D sin (α-φ) behind.

[0066] Resistance force F from the atmosphere is applied to each deployment film 14. Ｌ , F Ｒ are applied, the torque T generated on each deployment membrane 14 due to these resistance forces is Ｌｔ , T Ｒｔ are respectively T Ｌｔ = D x F Ｌｔ = D x F Ｌ cos(α+φ) T Ｒｔ = D x F Ｒｔ = D x F Ｒ cos(α-φ) and F Ｌ ≒F Ｒ In the case of T Ｒｔ >T Ｌｔ That is, in the state shown in FIG. 3B, a torque T applied backward to one of the deployable membranes 14 located relatively rearward in the traveling direction (X-axis direction) is Ｌｔ The torque T applied rearward to the other deployment membrane 14 located forward is Ｒｔ Therefore, the sum of these torques (T Ｒｔ -T Ｌｔ ) acts in a direction that tries to align the roll axis of the spacecraft 2 with the X axis, that is, acts as an attitude restoration torque.

[0067] Next, the attitude restoring torque when the spacecraft 2, 2' shown in Figs. 3A and 4A are tilted in different directions will be described with reference to Figs.

[0068] 5 and 6 show another example of a state in which resistance from the atmosphere is applied to the deployment membrane when the spacecraft 2 of the first embodiment or the spacecraft 2' of the reference example is flying with its roll axis tilted relative to the direction of travel. While Fig. 3(B) and Fig. 4(B) show the state in which the spacecraft 2, 2' are tilted around the pitch axis (Y axis) from the pitch axis (Y axis) direction, Fig. 5 and Fig. 6 show the state in which the spacecraft 2, 2' are tilted around the yaw axis (Z axis) from the yaw axis (Z axis) direction.

[0069] In Figure 5, the spacecraft 2 is tilted around the yaw axis (Z axis) at a slight angle ω with respect to the X-axis and Y-axis directions. In this case, the center of gravity G of the spacecraft 2 Ｓ On the other hand, the geometric center of gravity G of the deployment membrane 14 Ｗ is offset rearward with respect to the roll axis, the geometric center of gravity G Ｗ The position of the center of gravity G of Spacecraft 2 Ｓ Therefore, the resistance force F applied to the deployment film 14 from the atmosphere in the X-axis direction is a component perpendicular to the roll axis (F in FIG. 5). Ｓ As a result, the center of gravity G in the roll axis direction Ｓ , G Ｗ If the distance between them is d, the torque T is expressed by the following formula: Ｓ is the center of gravity G Ｓ Join the group around T Ｓ = d x F Ｓ = d × F sin ω This torque T Ｓ acts in the opposite direction to the tilt of the spacecraft 2 around the yaw axis (Z axis), and functions to correct the attitude of the spacecraft 2 so that the roll axis is parallel to the X axis. Ｓ acts as a posture restoring torque.

[0070] In contrast, the center of gravity G Ｓ , G Ｗ Considering the spacecraft 2' of the reference example in which no offset is set between the two (see FIG. 6), even if the spacecraft 2 tilts around the yaw axis (Z axis), the geometric center of gravity G of the deployment membrane 14 Ｗ The position of the center of gravity G of Spacecraft 2 Ｓ Since the position of the deployment film 14 coincides with the position of the deployment film 14, the resistance force F acting on the deployment film 14 does not have a component in the direction that rotates the spacecraft 2, and does not act to correct the inclination of the spacecraft 2.

[0071] 1, 3A, 3B, and 5, the geometric center of gravity G Ｗ is the center of gravity G of Spacecraft 2 Ｓ3B and 5, when the deployment membrane 14 is offset forward in the direction of travel (X axis), the center of gravity G of the spacecraft 2 is Ｓ The torque generated around the membrane 14, which is positioned relatively further rearward, is larger than the torque generated around the membrane 14, and the sum of the torques acts in a direction that strengthens the rotation of the spacecraft 2 (not shown). Therefore, when the center of gravity of the spacecraft and the geometric center of gravity of the membrane are offset in order to obtain the attitude restoring torque using the mechanisms shown in Figures 3(B) and 5, it is necessary to position the geometric center of gravity of the membrane rearward of the center of gravity of the spacecraft in the direction of travel (X-axis).

[0072] In actual operation of a spacecraft, the attitude of the spacecraft with respect to the direction of travel (X-axis) is not necessarily constant. Therefore, if the spacecraft is designed so that the geometric center of gravity of the deployment membrane is offset with respect to the roll axis from the center of gravity of the spacecraft as described above, it is possible that at one point during flight the geometric center of gravity of the deployment membrane is located behind the center of gravity of the spacecraft with respect to the direction of travel (X-axis), but at another point in time the geometric center of gravity of the deployment membrane is located ahead of the center of gravity of the spacecraft with respect to the direction of travel (X-axis).

[0073] The attitude restoring torque is calculated by the geometric center of gravity G of the deployable membrane 14. Ｗ and the center of gravity G of Spacecraft 2 Ｓ The moment of inertia increases as the offset between the centers of gravity (distance d in FIG. 5) increases, but on the other hand, the moment of inertia around the pitch axis and the yaw axis increases as the offset between the centers of gravity increases. If the moment of inertia around the pitch axis and the yaw axis becomes larger than the moment of inertia around the roll axis, the effect of gravity that tries to pull the spacecraft 2 down so that the roll axis is aligned with the direction of gravity increases. Therefore, it is preferable to design the spacecraft 2 so that the moment of inertia around the roll axis is larger than the moment of inertia around the other two axes (pitch axis and yaw axis) that intersect with the roll axis when the spacecraft 2 has deployed its deployment membrane 14.

[0074] In manufacturing the spacecraft 2 on which the above-described attitude restoration torque acts, a spacecraft main body 10 and a movement control device 12 having a deployment membrane 14 are prepared, and the movement control device 12 is attached to the main body 10 so that the geometric center of gravity of the deployment membrane 14 in the deployed state is offset in the roll axis direction from the center of gravity of the main body 10.

[0075] The above-described setting of the moment of inertia is effective under the condition that the atmosphere is sufficiently dense and the effect of the above-described attitude restoring torque obtained by the resistance of the atmosphere is sufficiently large compared to the action of gravity (for example, at a medium altitude or lower; approximately less than 600 km). However, at a high altitude, for example, above 600 km, the atmosphere is thin, and the resistance force that the spacecraft 2 receives from the atmosphere according to its speed becomes relatively small. Therefore, depending on the shape, weight, etc. of the spacecraft 2, the mechanical action exerted by gravity according to the moment of inertia becomes relatively large compared to the effect of the above-described attitude restoring torque.

[0076] If such conditions are assumed, the spacecraft 2 may behave undesirably due to the action of gravity, depending on the magnitude relationship of the moments of inertia around each axis.

[0077] Specifically, when the spacecraft 2 has a shape in which the moment of inertia around the pitch axis is smaller than the moment of inertia around the roll axis, if the attitude restoring torque due to atmospheric resistance as described above is not obtained sufficiently, it is assumed that the spacecraft 2 may rotate around the yaw axis due to the action of gravity. If this causes the pitch axis of the spacecraft 2 to swing significantly toward the direction of travel, the attitude restoring torque obtained by atmospheric resistance will become increasingly smaller, and there is a concern that maintaining the attitude will become increasingly difficult.

[0078] Therefore, research by the present inventors has revealed that under conditions such as high altitude where the available atmospheric resistance is relatively small, it is preferable that the moment of inertia around the pitch axis is sufficiently large. More specifically, it is preferable that the spacecraft 2 is set so that the moment of inertia around the pitch axis is larger than the moment of inertia around the roll axis.

[0079] Regarding the magnitude relationship between the moment of inertia around the roll axis and the moment of inertia around the yaw axis, it is preferable that the former is larger than the latter for the same reason as described above (if the moment of inertia around the yaw axis is larger than the moment of inertia around the roll axis, the effect of gravity that tries to pull the spacecraft 2 down so that the roll axis is aligned with the direction of gravity becomes larger).

[0080] To summarize the above, regarding the moment of inertia of the spacecraft 2, firstly, the moment of inertia around the roll axis is greater than the moment of inertia around the yaw axis (the moment of inertia around the roll axis is I ｘｘ , the moment of inertia around the yaw axis is I ｚｚ Then, I ｚｚ <I ｘｘ Furthermore, when flight is assumed to be performed at a relatively low altitude and in a relatively dense atmosphere, the moment of inertia around the roll axis is greater than the moment of inertia around the pitch axis and the moment of inertia around the yaw axis (the moment of inertia around the pitch axis is set to I ｙｙ Then, I ｚｚ <I ｘｘ And I ｙｙ <I ｘｘ It is preferable that the moment of inertia around the roll axis is greater than the moment of inertia around the yaw axis, and the moment of inertia around the pitch axis is greater than the moment of inertia around the roll axis (I ｚｚ <I ｘｘ <I ｙｙ It can be said that it is preferable that

[0081] The moment of inertia around each axis and the magnitude relationship between them can be adjusted, for example, by the offset amount in the roll axis direction of the geometric center of gravity of the deploying membrane 14 relative to the center of gravity of the main body 10, the angle of the deploying membrane 14 relative to each axis of the main body 10, the dimensions of the deploying membrane 14 extending around each axis, etc.

[0082] Second Embodiment In the first embodiment, the deployable membrane 14 is deployed along a plane perpendicular to the roll axis, and the geometric center of gravity G Ｗ is the center of gravity G of Spacecraft 2 ＳIn the second embodiment, a spacecraft 2 is described in which the deployable membrane 14 is configured to be deployable so as to form an inclined surface that is inclined at an acute angle with respect to the roll axis.

[0083] Figure 7 shows a spacecraft 2 according to a second embodiment. In the spacecraft 2 of the second embodiment shown in Figure 7, similar to the first embodiment shown in Figure 1, the deployment membranes 14 are deployed by the movement control device 12 so that the center of gravity of the spacecraft 2 coincides with the geometric center of gravity of the multiple deployment membranes 14 when viewed from the roll axis direction. However, the second embodiment differs from the first embodiment in that the movement control device 12 is provided with an angle adjustment mechanism that changes the angle of the deployment membranes 14.

[0084] 7, the movement control device 12 is equipped with a mechanism (angle adjustment mechanism) for raising and lowering the fixed members 18 constituting each deployment membrane 14 at the base end of the deployment membrane 14 relative to the main body 10, thereby adjusting the angle of the deployment membrane 14 with respect to the roll axis. As a result, the spacecraft 2 can deploy the deployment membrane 14 along a plane perpendicular to the roll axis, as in the first embodiment shown in FIG. 1, for example, but can also deploy the membrane 14 in an attitude inclined at an acute angle with respect to the roll axis, as shown in FIG.

[0085] The mechanism by which an attitude restoring torque is generated in the tilted deployment membrane 14 in the spacecraft 2 of the second embodiment will be described with reference to Fig. 8. Fig. 8 is a diagram schematically showing the resistance force applied from the atmosphere to the deployment membrane 14 when the roll axis of the spacecraft 2 is tilted with respect to the direction of travel (X axis) in the case where the spacecraft 2 of the second embodiment is flying with the two deployment membranes 14 deployed so that the line segment connecting their centers of gravity is along the yaw axis (Z axis).

[0086] For simplicity, the following description will be given taking as an example a case where the movement control device 12 deploys the two deployable membranes 14 rearward with respect to the roll axis relative to the center of gravity of the spacecraft 2 so that they form a plane passing through the center of gravity of the spacecraft 2 (in this case, the positions of the geometric centers of gravity of the multiple deployable membranes 14 are offset rearward with respect to the roll axis relative to the center of gravity of the spacecraft 2, as in the first embodiment).

[0087] The angle that the deployment membrane 14 makes with respect to the roll axis is assumed to be θ (0°<θ<90°). The two deployment membranes 14 are deployed symmetrically, each at an angle θ with respect to the roll axis, and the spacecraft 2 is flying with the roll axis parallel to the X axis. As shown in Figure 8, the spacecraft 2 tilts by a small angle φ around the pitch axis (Y axis) and assumes a posture in which one deployment membrane 14 (left side in the figure) is more forward than the other deployment membrane 14 (right side in the figure). In this case, the drag force that one deployment membrane 14 receives from the atmosphere in a direction perpendicular to the membrane surface 16 is F Ｌ cos(90°-θ+φ), and the drag force that the other deployment film 14 receives from the atmosphere in a direction perpendicular to the film surface 16 is F Ｒ Therefore, if the distance between the center of gravity of the spacecraft 2 and the center of gravity of each deployable membrane 14 is D, the torque T generated around the center of gravity of the spacecraft 2 with respect to each deployable membrane 14 is Ｌｔ , T Ｒｔ are respectively T Ｌｔ = D x F Ｌｔ = D x F Ｌ cos(90°-θ+φ) T Ｌｔ = D x F Ｒｔ = D x F Ｒ cos(90°-θ-φ).

[0088] In the figure, when the left and right deploying films 14 are symmetrical with respect to the roll axis, the ratio of the projected areas of the left and right deploying films 14 on a plane perpendicular to the direction of collision of the air molecules (X axis) is sin(θ-φ):sin(θ+φ). Ｌ , F Ｒ is proportional to the ratio of the projected areas, so F Ｌ , F Ｒ = sin(θ-φ): sin(θ+φ). Therefore, F Ｌ <F Ｒ and T Ｒｔ >T Ｌｔ The sum of these torques (T Ｒｔ -T Ｌｔ ) acts in a direction that tries to align the roll axis of the spacecraft 2 with the X axis, that is, acts as an attitude restoration torque.

[0089] 7, the spacecraft 2 of the second embodiment is equipped with a movement control device 12 configured to be able to control the tilt angle of the deployment membranes 14, and therefore it is possible to control the magnitude of the attitude restoring torque by adjusting the tilt angle of each deployment membrane 14. For example, the following mechanism can be used to control the attitude restoring torque.

[0090] 9 is a perspective view showing the configuration of the movement control device 12 provided in the spacecraft 2 of the second embodiment. As shown here, the movement control device 12 is provided with a rotation mechanism 24 for raising and lowering the deployment membrane 14. The rotation mechanism 24 is provided with a rotation shaft 44 oriented perpendicular to the roll axis, and the base end of the deployment membrane 14 is attached to the rotation shaft 44 so that the deployment membrane 14 can raise and lower around the rotation shaft 44. The movement control device 12 also has a sensor 40, which is a torque sensor that detects the torque acting on the deployment membrane 14.

[0091] For example, a two-axis force gauge can be used as the sensor 40. However, there is no limitation on the type of sensor 40 as long as it can detect the magnitude of the atmospheric resistance acting on the deployment membrane 14.

[0092] For example, the sensor 40 is attached to the housing 20 that houses the deployment membrane 14. When air molecules 34 collide with the deployment membrane 14 and atmospheric resistance acts, a torque acts on the deployment membrane 14, and the sensor 40 detects the torque transmitted from the deployment membrane 14 via the housing 20. However, as long as the sensor 40 can detect the torque acting on the deployment membrane 14, there are no restrictions on where the sensor 40 is attached. For example, the sensor 40 may be attached directly to the deployment membrane 14 and housed in the housing 20 together with the deployment membrane 14.

[0093] The rotation mechanism 24 rotates the deployment membrane 14 in response to atmospheric resistance acting on the deployment membrane 14. The rotation mechanism 24 can arbitrarily set the angle of the deployment membrane 14 relative to a rotation axis 44. Specifically, the rotation mechanism 24 is equipped with a rotation motor 42, and the deployment membrane 14, the housing portion 20, and the sensor 40 are connected to the output shaft of the rotation motor 42. The rotation motor 42 may be a servo motor, a stepping motor, or the like, which allows the rotation angle of the output shaft to be arbitrarily set. When the rotation motor 42 is operated, the deployment membrane 14, the housing portion 20, and the sensor 40 rotate around the rotation axis 44.

[0094] The movement control device 12 also includes an electronic computer 46. The electronic computer 46 corresponds to a control unit (controller, controller) that controls the movement control device 12. Specifically, the electronic computer 46 is connected to the components that make up the movement control device 12 (the deployment membrane 14, the sensor 40, the rotary motor 42, etc.), and controls the operation of the movement control device 12 by outputting control signals to each component.

[0095] For example, the electronic calculator 46 is configured by a computer. In this case, the electronic calculator 46 includes a processor such as a CPU (Central Processing Unit) that executes calculations and other processes necessary for controlling the movement control device 12, and memory such as a ROM (Read Only Memory) or RAM (Random Access Memory) that stores various information (data, programs, etc.) used to control the movement control device 12. The electronic calculator 46 executes programs stored in the memory to control the deployment and retraction of the deployable membrane 14, torque detection by the sensor 40, rotation of the deployable membrane 14 by the rotary motor 42, etc.

[0096] The information detected by the sensor 40 is input to the electronic computer 46. The electronic computer 46 controls the rotation motor 42 based on the information detected by the sensor 40 and sets the rotation angle of the deployment membrane 14. In this way, the rotation angle of the deployment membrane 14 is actively adjusted.

[0097] It should be noted that the movement control device 12 can be controlled by an electronic computer mounted on the main body 10 (see FIG. 1) of the spacecraft 2 instead of the electronic computer 46. In this case, the electronic computer 46 is omitted, and each component of the movement control device 12 is connected to the electronic computer provided on the main body 10. Then, the electronic computer provided on the main body 10 outputs control signals to each component of the main body 10 and the movement control device 12, thereby controlling the operations of the main body 10 and the movement control device 12.

[0098] The movement control device 12 also includes a battery 48 that supplies power to the components of the movement control device 12. A lithium-ion secondary battery, a fuel cell, or the like can be used as the battery 48. For example, the battery 48 supplies power to the sensor 40, the rotary motor 42, the computer 46, and the like.

[0099] The mobile control device 12 may be powered by another battery provided external to the mobile control device 12. For example, a battery provided in the main body 10 (see FIG. 1 ) of the spacecraft 2 or a battery provided independently of the main body 10 and the mobile control device 12 may supply power to the mobile control device 12. However, if the mobile control device 12 is equipped with a battery 48, the mobile control device 12 can be operated independently of the main body 10 (see FIG. 1 ) of the spacecraft 2 even when the battery provided external to the mobile control device 12 runs out of power or when it is difficult to supply power to the mobile control device 12 from an external source. The mobile control device 12 may also be powered by both the battery 48 and a battery provided external to the mobile control device 12.

[0100] Furthermore, the movement control device 12 includes a power generator 50 that generates power to operate the components of the movement control device 12. For example, the power generator 50 is a solar cell, and includes a photoelectric conversion element that converts light energy such as sunlight into electrical energy. However, there are no limitations on the type of power generator 50.

[0101] The generator 50 is appropriately installed in a position on the movement control device 12 that is suitable for generating power. For example, the generator 50 may be installed in the storage unit 20, the sensor 40, or the rotation mechanism 24. In the example shown here, the generator 50 is displayed near the movement control device 12 on the main body 10. Furthermore, if the movement control device 12 has a housing (not shown) that covers the storage unit 20, the sensor 40, and the rotation motor 42, one or more generators 50 may be installed on the outer circumferential surface of the housing. Furthermore, the generator 50 may also be attached to the deployment membrane 14.

[0102] Each of the power generators 50 generates electricity from energy such as sunlight and supplies it to other components of the movement control device 12. The electricity generated by the power generators 50 may be used to charge the battery 48, or to operate the sensor 40, the rotary motor 42, and the computer 46. However, if the batteries installed in the spacecraft 2 can provide all the electricity necessary to operate the movement control device 12, the power generators 50 may be omitted, and the batteries 48 installed in the movement control device 12 may also be omitted.

[0103] The mobile control device 12 may also include a communication unit (not shown) for performing information communication. The communication unit is a communication module for transmitting and receiving information to and from any communication device external to the mobile control device 12, and includes an antenna, a communication circuit, etc. By including the communication unit, the mobile control device 12 can communicate with the main body 10 of the spacecraft 2, other spacecraft, communication devices on the ground, etc.

[0104] As the spacecraft 2 travels along the X-axis, atmospheric molecules 34 collide with the membrane surface 16 of the deployable membrane 14 from the front to the rear in the direction of travel, applying an external force to the deployable membrane 14. The force acting on the deployable membrane 14 is then successively detected by the sensor 40, and the detected values ​​are input to the electronic computer 46. In this way, the force acting on the deployable membrane 14 is monitored.

[0105] Furthermore, the electronic computer 46 outputs a control signal to the rotation motor 42 based on the force detected by the sensor 40, thereby controlling the rotation motor 42. Then, based on the control signal input from the electronic computer 46, the rotation motor 42 sets the rotation angle of the deployable membrane 14 to an angle corresponding to the force detected by the sensor 40.

[0106] In particular, when the sensor 40 is a torque sensor, the torque acting on the deployable membrane 14 due to atmospheric resistance is detected by the sensor 40, and the torque of the deployable membrane 14 detected by the sensor 40 is input to the electronic computer 46. The electronic computer 46 then controls the rotation motor 42 based on the torque detected by the sensor 40, and the rotation motor 42 sets the rotation angle of the deployable membrane 14 to an angle corresponding to the torque detected by the sensor 40.

[0107] For example, the rotation mechanism 24 rotates the deployment membrane 14 in a direction that increases or decreases the atmospheric resistance acting on the deployment membrane 14. Specifically, as the deployment membrane 14 approaches a state in which it is positioned perpendicular to the direction of travel of the spacecraft 2 (the X-axis direction), the projected area of ​​the deployment membrane 14 in the direction perpendicular to the X-axis (the YZ plane) becomes larger, and the atmospheric resistance acting on the deployment membrane 14 increases. Therefore, the electronic computer 46 sequentially controls the rotation motor 42 so that the deployment membrane 14 rotates in a direction that increases the torque detected by the sensor 40.

[0108] Conversely, the closer the deployment membrane 14 is to being arranged parallel to the direction of travel (X-axis direction) of the spacecraft 2, the smaller the projected area of ​​the deployment membrane 14 in the direction perpendicular to the X-axis (YZ plane) becomes, and the less atmospheric resistance acts on the deployment membrane 14. Therefore, the electronic computer 46 sequentially controls the rotary motor 42 so that the deployment membrane 14 rotates in a direction that reduces the torque detected by the sensor 40.

[0109] In this way, by adjusting the rotation angle of the deployment membrane 14, it is possible to adjust the angle of the deployment membrane 14 with respect to the direction of travel of the spacecraft 2 and adjust the atmospheric resistance acting on the deployment membrane 14. This makes it possible to preferably perform operations such as correcting the attitude of the spacecraft 2 or releasing the spacecraft 2 from the circular orbit 6 (see FIG. 1).

[0110] As described above, the movement control device 12 according to the second embodiment includes the rotation mechanism 24 that can arbitrarily set the angle of the membrane surface 16 of the deployment membrane 14 relative to the roll axis. This makes it possible to adjust the rotation angle of the deployment membrane 14 so that a predetermined amount of atmospheric resistance acts on the deployment membrane 14 even if the attitude of the spacecraft 2 fluctuates, thereby making it possible to appropriately control the attitude and movement of the spacecraft 2.

[0111] 7 and 8, when the deployment membrane 14 is deployed so as to face forward from the spacecraft 2 with respect to the roll axis (when θ > 90° in FIGS. 7 and 8), the torque generated in the deployment membrane 14 located relatively rearward with respect to the X axis becomes larger than the torque generated around the center of gravity of the spacecraft 2 in the deployment membrane 14 located relatively forward with respect to the X axis, and the sum of the torques acts in a direction that strengthens the rotation of the spacecraft 2 (not shown). Therefore, when the deployment membrane 14 is deployed obliquely with respect to the roll axis in order to obtain attitude restoring torque using the mechanism shown in FIG. 7, the angle θ of the deployment membrane 14 needs to be an acute angle with respect to the rear of the roll axis.

[0112] In actual operation of a spacecraft, the attitude of the spacecraft with respect to the direction of travel (X-axis) is not necessarily constant. Therefore, if a spacecraft is designed so that the deployment membrane has an inclined surface with respect to the roll axis of the spacecraft as described above, at one point during flight, the surface of the deployment membrane may assume an attitude with respect to the direction of travel (X-axis) that is inclined at an acute angle to the rear of the roll axis, but at another point in time, the surface of the deployment membrane may assume an attitude with respect to the direction of travel (X-axis) that is inclined at an acute angle to the front of the roll axis (an obtuse angle to the rear of the roll axis).

[0113] Furthermore, the attitude restoring torque described above increases as the angle θ of the deployment membrane 14 relative to the roll axis decreases. However, as the angle θ decreases, the deployment membrane 14 protrudes further rearward and the diameter of the deployment membrane 14 around the roll axis decreases, resulting in a larger moment of inertia around the pitch axis and the yaw axis and a smaller moment of inertia around the roll axis. If the moments of inertia around the pitch axis and the yaw axis are larger than the moment of inertia around the roll axis, the gravitational force that tries to pull the spacecraft 2 down so that the roll axis is aligned with the direction of gravity increases. Therefore, when deploying the deployment membrane 14, it is preferable to design the spacecraft 2 or adjust the angle of the deployment membrane 14 so that the moment of inertia around the roll axis is larger than the moments of inertia around the other two axes (pitch axis and yaw axis) that intersect with the roll axis.

[0114] Regarding the setting of the moment of inertia, it has been explained in the above description of the first embodiment that it is preferable for the moment of inertia around the pitch axis to be sufficiently large because the attitude restoring torque due to atmospheric resistance becomes small under conditions such as high altitude. This point also applies to the spacecraft 2 of the second embodiment.

[0115] That is, regarding the moment of inertia of the spacecraft 2, first, the moment of inertia around the roll axis is greater than the moment of inertia around the yaw axis (I ｚｚ <I ｘｘ Furthermore, when flight is assumed to be performed at a relatively low altitude and in a relatively dense atmosphere, it is preferable that the moment of inertia around the roll axis is greater than the moment of inertia around the pitch axis and the moment of inertia around the yaw axis (I ｚｚ <I ｘｘ And I ｙｙ <I ｘｘ When flight at a relatively high altitude and in a relatively thin atmosphere is also assumed, it is preferable that the moment of inertia around the roll axis is larger than the moment of inertia around the yaw axis, and the moment of inertia around the pitch axis is larger than the moment of inertia around the roll axis (I ｚｚ <I ｘｘ <I ｙｙ It is preferable that

[0116] In the case of the spacecraft 2 of the second embodiment, the angle of the deployment membrane 14 can be adjusted by the movement control device 12. This allows the offset amount of the geometric center of gravity of the deployment membrane 14 relative to the center of gravity of the main body 10 in the roll axis direction and the extension amount of the deployment membrane 14 around the roll axis and pitch axis to be manipulated, thereby adjusting the moment of inertia around these axes. ｘｘ <I ｙｙ In the area where the altitude is relatively low, ｙｙ <I ｘｘ In theory, it is also possible to operate the movement control device 12 so that

[0117] The relationship between the inclination of the surface of the deployed membrane 14 and the attitude restoring torque has been described above, but to generate the attitude restoring torque as described above, the deployed membrane 14 does not necessarily need to have an inclined surface extending rearward at an acute angle to the roll axis over the entire membrane surface 16; such an inclination may be provided partially. For example, part of the membrane surface 16 may be a surface perpendicular to the roll axis, while the rest has an inclined surface as described above. Furthermore, the membrane surface 16 may be a curved surface, or may have portions with surfaces with different inclination angles.

[0118] Here, a case where the spacecraft 2 of the second embodiment is tilted in a different direction (about the yaw axis) relative to the X axis will be described with reference to FIG.

[0119] 10 shows a state in which the spacecraft 2 of the second embodiment is tilted at a slight angle ω around the yaw axis with respect to the X axis during flight while deploying the two deployment membranes 14 so that the line segment connecting their centers of gravity is along the yaw axis (Z axis). In this case, if the cross section of the surface of the deployment membrane 14 along the XY plane (roll axis-pitch axis plane) is perpendicular to the roll axis as viewed from the yaw axis (Z axis), the resistance force F that the deployment membrane 14 receives from the atmosphere along the X axis direction acts on the deployment membrane 14 parallel to the roll axis.

[0120] Therefore, for example, as shown in Figure 8, the direction of the resistance force acting on the deployment membrane 14 does not differ between the left and right sides of the figure. However, if the center of gravity of the deployment membrane 14 is offset rearward from the center of gravity of the spacecraft 2 with respect to the roll axis, an attitude restoring torque acts to reduce the tilt of the spacecraft 2 around the yaw axis, based on a principle similar to that shown in Figure 5 (the component of the resistance force acting as this attitude restoring torque is shown as F in Figure 10). Ｓ (Indicated by the symbol ).

[0121] In manufacturing the spacecraft 2 on which the above-described attitude restoring torque acts, the spacecraft main body 10 and the movement control device 12 having the deployment membrane 14 are prepared, and the movement control device 12 is attached to the main body 10 so that the geometric center of gravity of the deployment membrane 14 in the deployed state is offset in the roll axis direction from the center of gravity of the main body 10. In this case, the movement control device 12 is attached to the main body 10 so that the deployment membrane 14 in the deployed state forms an acute angle with respect to the rear of the roll axis.

[0122] In addition, the spacecraft 2 of the first embodiment (the center of gravity G of the spacecraft 2) as described in FIG. 3(B) and FIG. Ｓ On the other hand, the geometric center of gravity G of the deployment membrane 14 Ｗ The attitude restoring torque in the spacecraft 2 of the second embodiment (where the position of the deployment membrane 14 is offset rearward in the X-axis direction) and the attitude restoring torque in the spacecraft 2 of the second embodiment described in FIG. 8 (where the surface of the deployment membrane 14 extends rearward in relation to the roll axis from the base end to the tip, and is inclined at an acute angle to the roll axis) are generated by different principles and can occur independently of each other.

[0123] Therefore, for example, in a spacecraft in which the geometric center of gravity of the deployment membrane is located forward of the center of gravity of the spacecraft in the roll axis direction (Configuration A), and the surface of the deployment membrane extends rearward with respect to the roll axis from the base end to the tip, while being inclined at an acute angle with respect to the roll axis (Configuration B), if the roll axis of the spacecraft is inclined with respect to the X-axis during flight, and the torque generated by Configuration A in a direction that tends to increase the inclination of the spacecraft is exceeded by the attitude restoring torque by Configuration B, it is possible to correct the attitude of the spacecraft.

[0124] Conversely, in a spacecraft in which the geometric center of gravity of the deployment membrane is located rearward of the center of gravity of the spacecraft in the roll axis direction (Configuration C), and the surface of the deployment membrane extends forward with respect to the roll axis from the base end to the tip, while forming an inclined surface with respect to the roll axis (Configuration D), if the attitude restoring torque by Configuration C exceeds the torque that tends to increase the tilt of the spacecraft during flight, the attitude of the spacecraft can be corrected.

[0125] 11 is a perspective view showing the configuration of a spacecraft 2 according to a third embodiment. In the spacecraft 2 of the third embodiment, as in the second embodiment, the geometric center of gravity of the deployable membrane 14 is offset from the center of gravity of the spacecraft 2 with respect to the roll axis, and the movement control device 12 is attached to the main body 10 so that the deployable membrane 14 is deployed with an inclined surface with respect to the roll axis. Furthermore, in the spacecraft 2 of the third embodiment, each deployable membrane 14 can be rotated about the axis of the deployable membrane 14 (the axis extending along the plane of the deployable membrane 14 and radially outward from the roll axis of the spacecraft 2) by a rotation mechanism 24 provided in the movement control device 12.

[0126] The mechanism of the rotation mechanism 24 is generally similar to that of the rotation mechanism 24 in the second embodiment (see FIG. 9 ), but differs in the orientation of the rotation axis 44. That is, in the spacecraft 2 of the third embodiment, the orientation of the rotation axis 44 coincides with the axis of the deployment membrane 14, and the deployment membrane 14 rotates about the axis inclined at an acute angle with respect to the roll axis of the spacecraft 2.

[0127] In addition, the spacecraft 2 of the third embodiment is equipped with a sensor 40 that detects the torque acting on the deployment membrane 14, a rotary motor 42 that rotates the deployment membrane 14, a computer 46 that controls the movement control device 12, a battery 48 that supplies power to the components of the movement control device 12, a power generator 50, a communication unit, etc. (see FIG. 9 ). The torque acting on the deployment membrane 14 due to atmospheric resistance is detected by the sensor 40, and the angle of the deployment membrane 14 can be adjusted based on this.

[0128] (Fourth embodiment) Fig. 12 is a perspective view showing the configuration of a spacecraft 2 according to a fourth embodiment, and Fig. 13 is a side view showing the configuration of the spacecraft 2 of Fig. 12. The spacecraft 2 of this fourth embodiment comprises a cylindrical main body 10 and a movement control device 12 having two deployment membranes 14.

[0129] The central axis of the cylindrical body 10 extends longer than the diameter. That is, the longitudinal direction of the body 10 coincides with the axial direction. The body 10 also has a mass that is sufficiently larger than the mass of the deployment membrane 14.

[0130] Such a spacecraft 2 tries to assume a posture in which the longitudinal direction of the main body 10 is aligned with the direction of gravity due to the action of gravity. That is, in the spacecraft 2 of the fourth embodiment, the direction along the longitudinal direction of the columnar main body 10 becomes the yaw axis. Also, regarding the moment of inertia, the moment of inertia around the yaw axis (I ｚｚ ) is the moment of inertia (I) about the other two axes (roll and pitch axes). ｘｘ , I ｙｙ ) is smaller than the mass of the body 10 extending along the yaw axis. ｘｘ and I ｙｙ In other words, a member such as the deployable membrane 14, which is lighter than the main body 10, extends around the yaw axis, thereby ｚｚ Even if there is a slight increase in I ｚｚ Ga I ｘｘ and I ｙｙ It does not come close to exceeding this.

[0131] The spacecraft 2 of the fourth embodiment includes two deployable membranes 14. As shown in Figures 12 and 13, the base end of each deployable membrane 14 is attached to a position near the rear of the side surface of the cylindrical spacecraft 2, and extends obliquely rearward from there.

[0132] In a front view (a field of view looking at the spacecraft 2 from the front along the roll axis), the two deployable membranes 14 are arranged to protrude to the left and right from the main body 10. The membrane surfaces 16 of the two deployable membranes 14 form planes that are roughly parallel to the direction of gravity (Z-axis direction, yaw-axis direction) and extend rearward from the main body 10 while forming oblique angles with the XZ plane (a vertical plane including the roll axis and yaw axis) and the YZ plane (a vertical plane including the pitch axis and yaw axis). Thus, each membrane surface 16 forms a plane that intersects obliquely with the traveling direction of the spacecraft 2 (the direction along the roll axis) and is subjected to atmospheric resistance associated with the flight of the spacecraft 2.

[0133] The geometric center of gravity G of the two deployed membranes 14 ｗ is located at a position offset rearward with respect to the roll axis with respect to the center of gravity of the main body 10. As a result, the geometric center of gravity G ｗ is the center of gravity G of the entire spacecraft 2 ｓ It is also located rearward with respect to the roll axis.

[0134] The configuration of the deployment membrane 14 shown here is one example, and the number and shape of the deployment membrane 14, the attachment position and angle relative to the main body 10, etc. can be set appropriately in ways other than those shown in the example.

[0135] As for the setting of the moment of inertia, as in the first and second embodiments, I ｚｚ <I ｘｘ However, in the third embodiment, due to the shape and weight of the main body 10, ｚｚ <I ｘｘ At a relatively low altitude and in a relatively dense atmosphere, ｚｚ <I ｘｘ And I ｙｙ <I ｘｘ At a relatively high altitude and in a relatively thin atmosphere, it is preferable that ｚｚ <I ｘｘ <I ｙｙ It is preferable that:

[0136] The moment of inertia around each axis and the magnitude relationship between them can be adjusted, for example, by the offset amount in the roll axis direction of the geometric center of gravity of the deploying membrane 14 relative to the center of gravity of the main body 10, the angle of the deploying membrane 14 relative to each axis of the main body 10, the dimensions of the deploying membrane 14 extending around each axis, etc.

[0137] Furthermore, the movement control device 12 of the spacecraft 2 may be provided with a rotation mechanism 24 similar to that of the second and third embodiments, for example, so that the angle of the deployment membrane 14 relative to the main body 10 can be changed.

[0138] Fifth Embodiment Fig. 14 is a perspective view showing the configuration of a spacecraft 2 according to a fifth embodiment, and Fig. 15 is a side view showing the configuration of the spacecraft 2 of Fig. 14. The spacecraft 2 of this fifth embodiment comprises a cylindrical main body 10 and a movement control device 12 having two deployment membranes 14, similar to the spacecraft 2 of the fourth embodiment (see Figs. 12 and 13).

[0139] The main body 10 has a mass sufficiently large compared to the mass of the deployment membranes 14, and extends along the yaw axis with the central axis of the cylindrical shape as the longitudinal direction. The base ends of the two deployment membranes 14 are attached near the upper end of the main body 10.

[0140] In a front view (a field of view when the spacecraft 2 is viewed from the front along the roll axis), the two deployable membranes 14 are arranged so as to protrude downward to the left and right from near the upper end of the main body 10. The membrane surfaces 16 of the two deployable membranes 14 form planes that are roughly parallel to the pitch axis direction and extend rearward and downward from the main body 10 while forming oblique angles with the XY plane (a vertical plane including the roll axis and pitch axis) and the YZ plane (a vertical plane including the pitch axis and yaw axis). Thus, each membrane surface 16 forms a plane that intersects obliquely with the traveling direction of the spacecraft 2 (the direction along the roll axis) and is subjected to atmospheric resistance associated with the flight of the spacecraft 2.

[0141] The geometric center of gravity G of the two deployed membranes 14 ｗ is located at a position offset rearward with respect to the roll axis with respect to the center of gravity of the main body 10. As a result, the geometric center of gravity G ｗ is the center of gravity G of the entire spacecraft 2 ｓ It is also located rearward with respect to the roll axis.

[0142] In this way, the mounting position and angle of the deployment membrane 14 can be freely set according to the structure of the spacecraft 2, etc.

[0143] As described above, each embodiment provides a spacecraft 2 in which the geometric center of gravity of the deployable membrane 14 is offset with respect to the center of gravity of the spacecraft 2 with respect to the roll axis, and a spacecraft 2 in which the deployable membrane 14 in the deployed state has an inclined surface that forms an acute angle with the roll axis. As a result, when the attitude of the spacecraft 2 fluctuates, an attitude restoring torque is applied to correct the attitude fluctuation, making it possible to appropriately control the attitude and movement of the spacecraft 2.

[0144] The configurations, methods, etc. according to the above-described embodiments can be appropriately modified and implemented without departing from the scope of the present invention. Furthermore, the configurations, methods, etc. according to the present embodiments can be appropriately combined with other embodiments.

[0145] 2 spacecraft 4 celestial body 6 orbit 8 release orbit 10 main body 12 movement control device 14 deployment membrane (resistive membrane) 16 membrane surface 18 fixing member 20 storage section 24 rotation mechanism 34 atmospheric molecules 40 sensor 42 rotation motor 44 rotation shaft 46 electronic computer 48 battery 50 power generator

Claims

1. The spacecraft itself, The system comprises a deployable membrane that is deployed to exert atmospheric resistance, and a movement control device attached to the body of the spacecraft, which has an angle adjustment mechanism capable of adjusting the angle between the deployed membrane and the roll axis of the spacecraft in its deployed state. The aforementioned movement control device is attached to the spacecraft body such that the geometric center of gravity of the deployed membrane in its deployed state is offset with respect to the roll axis direction of the spacecraft body relative to the center of gravity of the spacecraft body. A spacecraft in which the moment of inertia around the roll axis is greater than the moment of inertia around the yaw axis.

2. The spacecraft according to claim 1, wherein the deployed film has an inclined surface that forms an acute angle with respect to the roll axis when deployed.

3. The spacecraft according to claim 1, wherein the spacecraft body or the movement control device has a computer that controls the operation of the angle adjustment mechanism.

4. The spacecraft body and A movement control device attached to the spacecraft body comprises a fixed member that can be maintained in a columnar shape, a pair of membrane surfaces fixed to the fixed member, and a deployable membrane that is deployed so as to exert atmospheric resistance, The aforementioned movement control device is attached to the spacecraft body such that the geometric center of gravity of the deployed membrane in its deployed state is offset with respect to the roll axis direction of the spacecraft body relative to the center of gravity of the spacecraft body. A spacecraft in which the moment of inertia around the roll axis is greater than the moment of inertia around the yaw axis.

5. The spacecraft according to claim 4, wherein the deployed film has an inclined surface that forms an acute angle with respect to the roll axis when deployed.

6. The spacecraft according to claim 4, wherein the movement control device further comprises an angle adjustment mechanism capable of adjusting the angle between the deployed membrane in its deployed state and the roll axis of the spacecraft.

7. The spacecraft according to claim 6, wherein the spacecraft body or the movement control device has a computer that controls the operation of the angle adjustment mechanism.

8. The spacecraft according to any one of claims 1 to 7, wherein the mobile control device further comprises one or both of a solar cell and a battery.

9. The spacecraft according to any one of claims 1 to 7, wherein the movement control device causes the spacecraft body to leave orbit by deploying the deployment membrane.

10. A spacecraft body, a deployable membrane that is deployed to exert atmospheric resistance, and a movement control device having an angle adjustment mechanism that can adjust the angle between the deployed membrane and the roll axis of the spacecraft are prepared. A method for manufacturing a spacecraft, comprising attaching the movement control device to the spacecraft body such that the geometric center of gravity of the deployed film in its deployed state is offset with respect to the center of gravity of the spacecraft body in the direction of the roll axis.

11. The spacecraft body, A moving control device is prepared, which has a deployable membrane that is deployed so as to be subject to atmospheric resistance, comprising a fixed member that can be maintained in a columnar shape and a pair of membrane surfaces fixed to the fixed member, A method for manufacturing a spacecraft, comprising attaching the movement control device to the spacecraft body such that the geometric center of gravity of the deployed film in its deployed state is offset with respect to the center of gravity of the spacecraft body in the direction of the roll axis.

12. A method for manufacturing a spacecraft according to claim 10 or 11, wherein the movement control device is attached to the spacecraft body such that the deployed membrane forms an acute angle with respect to the rear of the roll axis when deployed.

13. The spacecraft itself, The system comprises a deployable membrane that is deployed to exert atmospheric resistance, and a movement control device attached to the body of the spacecraft, which has an angle adjustment mechanism capable of adjusting the angle between the deployed membrane and the roll axis of the spacecraft in its deployed state. The aforementioned movement control device is attached to the spacecraft body such that the deployed membrane, when extended, has an inclined surface that forms an acute angle with respect to the roll axis. A spacecraft in which the moment of inertia around the roll axis is greater than the moment of inertia around the yaw axis.

14. The spacecraft body and A movement control device attached to the spacecraft body comprises a fixed member that can be maintained in a columnar shape, a pair of membrane surfaces fixed to the fixed member, and a deployable membrane that is deployed so as to exert atmospheric resistance, The aforementioned movement control device is attached to the spacecraft body such that the deployed membrane, when extended, has an inclined surface that forms an acute angle with respect to the roll axis. A spacecraft in which the moment of inertia around the roll axis is greater than the moment of inertia around the yaw axis.

15. A spacecraft body, a deployable membrane that is deployed to exert atmospheric resistance, and a movement control device having an angle adjustment mechanism that can adjust the angle between the deployed membrane and the roll axis of the spacecraft are prepared. A method for manufacturing a spacecraft, comprising attaching the movement control device to the spacecraft body such that the deployed membrane, when unfolded, has an inclined surface that forms an acute angle with respect to the roll axis.

16. The spacecraft body, A moving control device is prepared, which has a deployable membrane that is deployed so as to be subject to atmospheric resistance, comprising a fixed member that can be maintained in a columnar shape and a pair of membrane surfaces fixed to the fixed member, A method for manufacturing a spacecraft, comprising attaching the movement control device to the spacecraft body such that the deployed membrane, when unfolded, has an inclined surface that forms an acute angle with respect to the roll axis.