Spacecraft and manufacturing method thereof

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

[Technical Field]

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

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

[0003] If a spacecraft remains in orbit around the Earth after completing its intended mission, it may become space debris and hinder the operation of other spacecraft. Therefore, it is necessary to remove a spacecraft from orbit after its mission is complete. However, removing a spacecraft from orbit using propulsion mechanisms such as engines requires extensive control of the spacecraft's movement, which is time-consuming and costly. In addition, a spacecraft may not have enough energy remaining after its mission is complete to remove it from orbit.

[0004] Therefore, spacecraft are sometimes equipped with a movement control device (orbital de-orbiting device) to remove the spacecraft from orbit. Such movement control devices are also called PMD (Post Mission Disposal) devices, and they can operate independently of the spacecraft itself to remove the spacecraft from orbit.

[0005] For example, a motion control device is used that extends a tape-like conductive tether into space to decelerate a spacecraft (see Patent Document 1). When the conductive tether extended from the motion 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 orbit.

[0006] However, controlling a long conductive tether to extend correctly in a predetermined direction from a motion control device moving in orbit is highly difficult. Furthermore, in order to control the movement of a spacecraft with high precision using a conductive tether, the electron emitter that emits electrons at the end of the conductive tether must be constantly and stably operating. Therefore, a high level of reliability is required for the control of the motion control device, and power consumption is also high.

[0007] Therefore, a motion control device that uses atmospheric drag to decelerate a spacecraft has also been proposed (see Patent Document 2). This type of motion control device is equipped with a deployable membrane that is subject to atmospheric drag, and the membrane is deployed and spread out when the spacecraft de-orbits. As a result, atmospheric drag acts on the deployed membrane, decelerating the spacecraft and allowing it to de-orbit. [Prior art documents] [Patent Documents]

[0008] [Patent Document 1] Japanese Patent Publication No. 2021-115713 [Patent Document 2] Japanese Patent Publication No. 2022-143395 [Overview of the project] [Problems that the invention aims to solve]

[0009] By equipping a spacecraft with a movement control system, it becomes possible to decelerate the spacecraft and remove it from Earth's orbit, or to temporarily change the spacecraft's direction of travel or speed. Furthermore, by adopting a method of controlling the spacecraft's movement using atmospheric resistance acting on a deployable membrane, the movement control of the spacecraft is simplified compared to using a conductive tether.

[0010] In controlling the movement of a spacecraft, it is preferable for the spacecraft to maintain a constant attitude as much as possible. However, spacecraft in orbit do not always move in a constant attitude, and tilting or rotation may occur due to some external force. As a result, the movement control capability of the movement control system deteriorates, making it difficult to smoothly perform operations such as deorbiting the spacecraft.

[0011] Furthermore, when the movement control system deploys the deployment membrane, the mass characteristics of the entire spacecraft change. This makes it difficult for the spacecraft to maintain a constant attitude, and it may enter a state of movement accompanied by rotational motion (tumbling mode). When the spacecraft enters tumbling mode, controlling its movement by the movement control system becomes more difficult. As a result, it may become impossible to deorbit the spacecraft, 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 its landing site. In addition, it becomes difficult to control the spacecraft by temporarily changing its direction of travel or speed to avoid collision with space debris or other objects that are approaching it.

[0012] This invention has been made in view of the above problems, and aims to provide a spacecraft capable of appropriately controlling its attitude and a method for manufacturing the same. [Means for solving the problem]

[0014] According to one aspect of the present invention , u A spacecraft is provided, 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 as to exert atmospheric drag, wherein the 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 center of gravity of the spacecraft body with respect to the roll axis of the spacecraft, and the moment of inertia of the spacecraft around the roll axis is greater than the moment of inertia of the spacecraft around the yaw axis.

[0016] Preferably, the unfolded film has an inclined surface that forms an acute angle with respect to the roll axis.

[0017] Preferably, the movement control device further includes an angle adjustment mechanism capable of adjusting the angle between the deployed membrane and the roll axis of the spacecraft in its deployed state. Preferably, the unfolded film comprises a fixing member that can be maintained in a columnar shape, and a film surface fixed to a pair of the fixing members. The pair of fixing members to which the membrane surface is fixed are spaced further apart from each other as they move away from the spacecraft, and the membrane surface is formed to be wider as it moves away from the spacecraft body. ru.

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

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

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

[0021] According to another aspect of the present invention 、 A method for manufacturing a spacecraft is provided, which involves preparing a spacecraft body and a movement control device having a deployable membrane that is deployed to exert atmospheric drag, and attaching the movement control device to the spacecraft body such that the geometric center of gravity of the deployed membrane is offset with respect to the center of gravity of the spacecraft body in the direction of the roll axis. Preferably, the movement control device includes an angle adjustment mechanism capable of adjusting the angle between the deployed membrane and the roll axis of the spacecraft. Preferably, the deployed membrane comprises a fixing member that can be maintained in a columnar shape and a membrane surface fixed to a pair of the fixing members, wherein the pair of fixing members to which the membrane surface is fixed are separated from each other as they move away from the spacecraft, and the membrane surface is formed to be wider as it moves away from the spacecraft body.

[0022] Preferably, the movement control device is mounted on the spacecraft body such that the deployed membrane forms an acute angle with respect to the rear of the roll axis when deployed.

[0024] According to yet another aspect of the present invention 、 A spacecraft is provided, comprising a spacecraft body and a movement control device attached to the spacecraft body, the movement control device being attached to the spacecraft body such that the deployed membrane has an inclined surface that forms an acute angle with respect to the roll axis, wherein the spacecraft's moment of inertia around the roll axis is greater than its moment of inertia around the yaw axis. Preferably, the movement control device includes an angle adjustment mechanism capable of adjusting the angle between the deployed membrane and the roll axis of the spacecraft. Preferably, the deployed membrane comprises a fixing member that can be maintained in a columnar shape and a membrane surface fixed to a pair of the fixing members, wherein the pair of fixing members to which the membrane surface is fixed are separated from each other as they move away from the spacecraft, and the membrane surface is formed to be wider as it moves away from the spacecraft body.

[0026] According to yet another aspect of the present invention 、 A method for manufacturing a spacecraft is provided, which involves preparing a spacecraft body and a movement control device having a deployable membrane that is deployed to exert atmospheric drag, and attaching the movement control device to the spacecraft body such that the deployed membrane has an inclined surface that forms an acute angle with respect to the roll axis. Preferably, the movement control device includes an angle adjustment mechanism capable of adjusting the angle between the deployed membrane and the roll axis of the spacecraft. Preferably, the deployed membrane comprises a fixing member that can be maintained in a columnar shape and a membrane surface fixed to a pair of the fixing members, wherein the pair of fixing members to which the membrane surface is fixed are separated from each other as they move away from the spacecraft, and the membrane surface is formed to be wider as it moves away from the spacecraft body. [Effects of the Invention]

[0027] The present invention provides a spacecraft in which the geometric center of gravity of the deployable membrane is offset from the center of gravity of the spacecraft with respect to the roll axis, and a spacecraft in which the deployed membrane has an inclined surface that forms an acute angle with respect to the roll axis. This makes it possible to appropriately control the attitude and movement of the spacecraft by applying an attitude restoration torque to correct the attitude change when the attitude of the spacecraft changes. [Brief explanation of the drawing]

[0028] [Figure 1] Figure 1 is a perspective view showing a spacecraft according to the first embodiment. [Figure 2] Figure 2(A) is a front view showing the movement control device with the unfolded film contained, and Figure 2(B) is a front view showing the movement control device with the unfolded film deployed. [Figure 3] Figure 3(A) schematically shows an example of a state in which atmospheric drag is applied to the deployable membrane when the spacecraft of the first embodiment is flying in an attitude in which the roll axis is parallel to the direction of travel, and Figure 3(B) schematically shows an example of a state in which atmospheric drag is applied to the deployable membrane when the spacecraft of the first embodiment is flying in an attitude in which the roll axis is tilted with respect to the direction of travel. [Figure 4] Figure 4(A) schematically shows an example of atmospheric drag applied to the deployable membrane when the reference spacecraft is flying with its roll axis parallel to its direction of travel, and Figure 4(B) schematically shows an example of atmospheric drag applied to the deployable membrane when the reference spacecraft is flying with its roll axis tilted relative to its direction of travel. [Figure 5] Figure 5 schematically shows another example of a state in which atmospheric drag is applied to the deployed membrane when the spacecraft of the first embodiment is flying with its roll axis tilted relative to the direction of travel. [Figure 6] Figure 6 schematically shows another example of the situation in which atmospheric drag is applied to the deployed membrane when the reference example spacecraft is flying with its roll axis tilted relative to the direction of travel. [Figure 7]Figure 7 is a perspective view showing a spacecraft according to the second embodiment. [Figure 8] Figure 8 schematically shows an example of a state in which atmospheric drag is applied to the deployed membrane when the spacecraft of the second embodiment is flying with its roll axis tilted relative to the direction of travel. [Figure 9] Figure 9 is a perspective view showing the configuration of the mobile control device installed in the spacecraft of the second embodiment. [Figure 10] Figure 10 schematically shows another example of a state in which atmospheric drag is applied to the deployed membrane when the spacecraft of the second embodiment is flying with its roll axis tilted relative to the direction of travel. [Figure 11] Figure 11 is a perspective view showing a spacecraft according to the third embodiment. [Figure 12] Figure 12 is a perspective view showing a spacecraft according to the fourth embodiment. [Figure 13] Figure 13 is a side view of the spacecraft shown in Figure 12. [Figure 14] Figure 14 is a perspective view showing a spacecraft according to the fifth embodiment. [Figure 15] Figure 15 is a side view of the spacecraft shown in Figure 14. [Modes for carrying out the invention]

[0029] (First Embodiment) Hereinafter, an embodiment of one aspect of the present invention will be described with reference to the attached drawings. First, an example of the configuration of a spacecraft according to this embodiment will be described. Figure 1 is a perspective view showing the spacecraft 2.

[0030] Spacecraft 2 is a flying object that orbits a celestial body 4, such as Earth, along a circular orbit 6. For example, spacecraft 2 may be an artificial satellite, rocket, transport vehicle, space probe, or space station, and may be equipped with a propulsion mechanism such as a chemical propulsion engine or an electric propulsion engine. However, there are no restrictions on the type of spacecraft 2, and spacecraft 2 may be a flying object without a propulsion mechanism.

[0031] Orbit 6 is the travel path of spacecraft 2, set outside of celestial body 4. For example, orbit 6 is a roughly circular or elliptical path set so that the distance (altitude) from the surface (ground) of celestial body 4 remains roughly constant. The altitude of orbit 6 is set according to the type of spacecraft 2 and its operational purpose, for example, between 200 km and 1000 km.

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

[0033] Furthermore, spacecraft 2 has three mutually perpendicular axes—a roll axis, a pitch axis, and a yaw axis—that pass through its center of gravity. Spacecraft 2 is designed assuming that the roll axis, pitch axis, and yaw axis are positioned parallel to the X axis, Y axis, and Z axis, respectively. Therefore, the direction of the roll axis of spacecraft 2 corresponds to the direction assumed to be the direction of spacecraft 2's movement. When moving spacecraft 2 in orbit 6, the attitude of 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 the sake of clarity, the following explanation assumes that the roll, pitch, and yaw axes of spacecraft 2, positioned in orbit 6, are parallel to the X, Y, and Z axes, respectively. However, when spacecraft 2 actually moves along orbit 6, its attitude will fluctuate, and its roll, pitch, and yaw axes may not be parallel to the X, Y, and Z axes, respectively.

[0035] During the operation of Spacecraft 2, it will orbit celestial body 4 along orbital orbit 6, carrying out pre-specified missions (such as collecting and monitoring information on the ground and in space, conducting experiments in the space environment, communicating with the ground, and transporting goods). Once the mission is complete and Spacecraft 2's operation ends, Spacecraft 2 will leave orbital orbit 6 and enter a departure trajectory 8, gradually approaching celestial body 4 while orbiting it. After that, Spacecraft 2 will either re-enter the atmosphere and burn up, or reach celestial body 4 and be recovered. The time required for Spacecraft 2 to deorbit will vary depending on the type of Spacecraft 2, altitude, etc., and may range from several hours, weeks, months, years, or even more than 10 years.

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

[0037] The movement control device 12 includes multiple deployable membranes (resistance membranes) 14 that are deployed to exert atmospheric resistance. The multiple deployable membranes 14 are arranged at approximately equal angular intervals around the roll axis and are positioned symmetrically with respect to the roll axis of the spacecraft 2. For example, as shown in Figure 1, if the movement control device 12 includes two sets of deployable membranes 14, the two sets of deployable membranes 14 are attached to both ends of the main body 10 in the yaw axis direction, i.e., the upper and lower ends. This results in the two sets of deployable membranes 14 being positioned symmetrically at 180° intervals around the roll axis.

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

[0039] Furthermore, the arrangement of the deployable membrane 14 does not necessarily have to be symmetrical with respect to the roll axis, nor does it have 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 membrane 14 coincides with the position of the center of gravity of the entire spacecraft 2 when viewed from the direction of the roll axis.

[0040] In this specification, the geometric centroid of the deployable membrane 14 refers to "the position of the centroid of all deployable membranes 14 provided on the spacecraft 2." For example, if the spacecraft 2 is equipped with two deployable membranes 14 symmetrically around the main body 10, the geometric centroid of the deployable membranes 14 lies on a hypothetical line segment connecting the centroids of each deployable membrane 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 stretched state. The film constitutes a 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 ends of the membrane surface 16 are fixed to the fixing members 18.

[0042] The material, shape, dimensions, etc., of the film constituting the film surface 16 are not limited as long as the film surface 16 can withstand atmospheric resistance. For example, the film is constructed by covering the surface of the resin film with a conductive thin film. This prevents the resin film from degrading due to exposure to atomic oxygen (AO) present in the orbital 6. Furthermore, by forming a conductive thin film on the resin film and imparting conductivity to the film, damage to the film due to charging and discharging can be avoided.

[0043] The aforementioned film comprises a resin film (approximately 12.5 μm thick) 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 m². 2 It can be set to a certain extent. However, the material, shape, size, etc. of the film constituting the membrane surface 16 can be appropriately selected according to the weight, shape, size, altitude, etc. of the spacecraft 2.

[0044] The fixing member 18 is a member that can be maintained in a columnar shape and may be deformable or expandable. For example, a cylindrical member made of carbon fiber reinforced plastic (CFRP) can be used as the fixing member 18. The thickness of the fixing member 18 (difference between outer and inner diameters) 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 appropriately selected according to the shape, size, etc. of the membrane surface 16.

[0045] Furthermore, the fixing member 18 may be a cylindrical member (inflatable tube) that expands by filling its interior with gas and is fixed in a columnar shape. An inflatable tube is a deformable, hollow cylindrical member, and is constructed, for example, by rolling a film such as a resin film, metal film, or carbon film into a cylindrical shape.

[0046] If 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 in a columnar form.

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

[0048] The fixing member 18 may be an insulator or a conductor. If the fixing member 18 is conductive, 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 celestial body 4 to control the movement of spacecraft 2.

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

[0050] Figure 2(A) is a front view showing the movement control device 12 with the unfolded film 14 housed inside. The movement control device 12 includes a housing section 20 for housing the unfolded film 14. For example, the housing section 20 is a box-shaped container that houses the folded or rolled film surface 16 and fixing member 18. There are no restrictions on the shape and size of the housing section 20, as long as the unfolded film 14 can be housed inside it.

[0051] Figure 2(B) is a front view showing the movement control device 12 in the state where the unfolded film 14 is unfolded. When the movement control device 12 is activated, a pair of fixing members 18 are ejected from the housing section 20 so as to stretch in a predetermined direction due to the action of their own elastic energy, etc. For example, the pair of fixing members 18 stretch in a direction inclined with respect to the yaw axis (Z axis) such that their tips are separated from each other along the pitch axis (Y axis). When the shape of the fixing members 18 is fixed in a columnar shape, the film surface 16 attached to the fixing members 18 becomes taut, and the unfolded film 14 is unfolded.

[0052] The timing for removing spacecraft 2 from orbit 6 is set appropriately according to the content of the mission performed by spacecraft 2 and the lifespan of 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 spacecraft 2 or from the ground. Alternatively, the mobile control device 12 may deploy the deployable membrane 14 when the operating time or flight time of spacecraft 2 reaches a predetermined time limit. Furthermore, the mobile control device 12 may deploy the deployable membrane 14 when it receives a signal indicating that the operation of the main body 10 of spacecraft 2 has stopped.

[0053] Here, in particular, in the spacecraft 2 of this first embodiment, as shown in Figure 1, the overall geometric centroid of the deployed film 14 deployed around the roll axis (G in the figure) W The position of the center of gravity of spacecraft 2 (indicated by the symbol G in the figure) is as shown in Figure 1. SThe spacecraft 2 and the deployable membrane 14 are designed to be offset in the roll axis direction with respect to the (indicated by the sign) of the roll axis. This ensures that the attitude of the spacecraft 2 during flight is maintained with the roll axis aligned with the X axis. The effects of this attitude maintenance are explained below.

[0054] For the sake of explanation, we will use the term "G-force of spacecraft 2". S The explanation states that "the center of gravity of the entire spacecraft 2, including the deployable membrane 14, when the deployable membrane 14 is deployed" but if the center of gravity G S Even if we define it as "the center of gravity of the spacecraft body 10 excluding the deployable membrane 14," the effect is essentially the same in terms of achieving the effects described below. This is because, regardless of which definition is adopted, the geometric center of gravity of the deployable membrane 14 will be offset in the direction of the roll axis relative to the center of gravity of the entire spacecraft 2, including the deployable membrane 14.

[0055] Figure 3(A) schematically shows an example of a state in which atmospheric drag is applied to the deployable membrane when the spacecraft 2 of the first embodiment is flying in an attitude in which the direction of travel and the roll axis are parallel. Figure 4(A) similarly schematically shows an example of a state in which atmospheric drag is applied to the deployable membrane when the spacecraft is flying in an attitude in which the direction of travel and the roll axis are parallel. However, as a reference example for comparison with spacecraft 2 in Figure 3(A), spacecraft 2' is shown 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 coincide).

[0056] In Figures 3(A) and 4(A), two deployable membranes 14 are deployed symmetrically along the yaw axis around the roll axis of spacecraft 2(2') (orienting them to be 180° apart when viewed from the roll axis direction), and this state is viewed from a direction perpendicular to the roll axis and yaw axis (the line connecting the centers of gravity of the two deployable membranes 14) (the pitch axis (Y axis) direction).

[0057] In the spacecraft 2 shown in Figure 3(A), the center of gravity G of spacecraft 2 is the same as that of spacecraft 2 shown in Figure 1. S The geometric centroid G of the two unfolded films 14W are offset with respect to the roll axis, and the geometric center of gravity G W is the center of gravity G S is located behind the center of gravity G S of the spacecraft 2 with respect to the traveling direction of the spacecraft 2 along the roll axis (X axis). In contrast, in the spacecraft 2' shown in FIG. 4(A), the center of gravity G W of the spacecraft 2' and the geometric center of gravity G

[0058] of the two deployment membranes 14 are not offset and coincide with each other. During the flight of the spacecrafts 2 and 2', the deployment membranes 14 are deployed so as to form a plane intersecting the traveling direction (X-axis direction) of the spacecrafts 2 and 2'. Therefore, each deployment membrane 14 receives air resistance in the direction from the front to the rear with respect to the traveling direction. When the two deployment membranes 14 are deployed symmetrically with respect to the roll axis along the yaw axis, each deployment membrane 14 located on the left and right in the figure receives substantially equal air resistance F L , F R .

[0059] Here, in the spacecraft 2' of the reference example shown in FIG. 4(A), 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, considering the torque generated around the center of gravity G S of the spacecraft 2', the torques due to the air resistances F L , F R act in opposite directions and are equal. That is, if the distances from the center of gravity G W to the centers of gravity G L , G R of the respective deployment membranes 14 are each D, torques of D×F S , D×F L , D×F R act in opposite directions around the center of gravity G

[0060] In the spacecraft 2 of the first embodiment shown in FIG. 3(A), it is substantially the same. As shown in FIG. 3(A), the center of gravity G S of the spacecraft 2 and the geometric center of gravity G WWhen the two are offset with respect to the roll axis, the resistance force that each of the left and right deployed films 14 receives from the atmosphere along the X-axis is F L ,F R If that is the case, then the center of gravity G of spacecraft 2 S And the center of gravity G of each unfolded film 14 L ,G R The component F is perpendicular to the line segment connecting the two points. Lt ,F Rt However, this acts as a torque that attempts to rotate spacecraft 2 around the pitch axis (Y axis). Center of gravity G S From the center of gravity G L ,G R If the distance to each point is D, then the center of gravity G S Around D×F Lt , D×F Rt Since the torques act in opposite directions, these torques cancel each other out, and spacecraft 2 still does not rotate.

[0061] Next, let's consider the case where the roll axis of spacecraft 2,2' is tilted with respect to the direction of travel (X-axis direction). Figure 4(B) shows the center of gravity G. S ,G W This schematic diagram illustrates an example of atmospheric drag applied to the deployed membrane when spacecraft 2', a reference example where the two spacecraft are not offset from each other, is flying with its roll axis tilted relative to the direction of travel.

[0062] In Figure 4(B), spacecraft 2' is tilted at a slight angle φ with respect to the X-axis and Z-axis directions around the pitch axis (Y-axis). In this case, the drag force F experienced by one of the deployed membranes 14 from the atmosphere is L Torque D×F Lt And the other developing film 14 receives resistance force F from the atmosphere. R Torque D×F Rt These are equal to each other. Therefore, if an external force acts on the spacecraft 2' from the state shown in Figure 4(A), and the spacecraft 2' rotates and tilts by an angle φ as shown in Figure 4(B), the rotation of the spacecraft 2' is equal to the resistance force F that the deployed membrane 14 receives from the atmosphere. L ,F R It will not be stopped by that.

[0063] Here, the center of gravity G S ,G W Let's consider the case where the spacecraft 2 of the first embodiment, in which the components are offset from each other, is similarly tilted with respect to the X-axis. Figure 3(B) schematically shows an example of the state in which atmospheric drag is applied to the deployable membrane when the spacecraft 2 of the first embodiment is flying in an attitude in which the roll axis is 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 The center of gravity of spacecraft 2 S It is located further back than that.

[0064] Assuming that spacecraft 2 is tilted by an angle φ with respect to the X-axis and Z-axis directions around the pitch axis (Y-axis), as shown in Figure 3(B), the drag force F exerted on the deployed film 14 by the atmosphere L ,F R The center of gravity G of spacecraft 2 S Components that form around (resistance F) L ,F R Of these, the center of gravity G of spacecraft 2 S And the center of gravity G of each unfolded film 14 L ,G R (Component in the direction perpendicular to the line segment connecting and) F Lt ,F Rt They are not equal. Let's assume resistance force F L ,F R Even if they are equal to each other, these resistance forces F L ,F R Because the position where it is added is not symmetrical with respect to the X-axis, spacecraft 2 has a center of gravity G S The component F that rotates around the object. Lt ,F Rt They are different.

[0065] The two unfolded membranes 14 have a center of gravity G L ,G R The position of the center of gravity G of spacecraft 2 S As a result of being mounted to spacecraft 2 so as to be equally offset with respect to the roll axis, the center of gravity G of spacecraft 2 is as shown in Figure 3(A). S Center of gravity G of each unfolded film 14 as seen from L ,G RThe positions are each located rearward by an amount corresponding to an angle α with respect to the yaw axis (center of gravity G S and center of gravity G L , G R Let the distances between them be D respectively. Then, the center of gravity G L , G R is located rearward by Dsinα with respect to the center of gravity G S with respect to the roll axis). From this state, as shown in Fig. 3(B), assume that the spacecraft 2 is tilted by a small angle φ about the pitch axis (Y axis). Then, the position of the center of gravity G L of one of the deployment membranes 14 is located rearward by Dsin(α + φ) with respect to the X axis as viewed from the center of gravity G S of the spacecraft 2, and the position of the center of gravity G R of the other deployment membrane 14 is located rearward by Dsin(α - φ) as viewed from the center of gravity G S of the spacecraft 2.

[0066] Assume that resistive forces F L , F R are applied to each of the deployment membranes 14. Then, the torques T Lt , T Rt generated in each of the deployment membranes 14 by these resistive forces are respectively T Lt = D×F Lt = D×F L cos(α + φ) T Rt = D×F Rt = D×F R cos(α - φ) and when F L ≈ F R , T Rt > T Lt . That is, in the state shown in Fig. 3(B), the torque T Lt acting rearward on one of the deployment membranes 14 located relatively rearward with respect to the traveling direction (X-axis direction) is smaller than the torque T Rt acting rearward on the other deployment membrane 14 located forward. Therefore, the sum of these torques (T Rt - T Lt ) acts in the direction to align the roll axis of the spacecraft 2 with the X axis, that is, as a posture restoration torque.

[0067] Next, the attitude restoration torque when spacecraft 2,2' shown in Figures 3(A) and 4(A) are tilted in a different direction will be explained with reference to Figures 5 and 6.

[0068] Figures 5 and 6 schematically show another example of atmospheric drag applied to the deployed 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. In Figures 3(B) and 4(B), the spacecraft 2 and 2' are shown tilted around the pitch axis (Y axis) from the direction of the pitch axis (Y axis), while in Figures 5 and 6, the spacecraft 2 and 2' are shown tilted around the yaw axis (Z axis) from the direction of the yaw axis (Z axis).

[0069] In Figure 5, spacecraft 2 is tilted at a slight angle ω around the yaw axis (Z axis) with respect to the X and Y axes. In this case, the center of gravity G of spacecraft 2 is S In contrast, the geometric centroid G of the unfolded film 14 W Due to its rearward offset position with respect to the roll axis, the geometric centroid G of the unfolded film 14 W The position is the center of gravity G of spacecraft 2. S In contrast, it shifts in the Y-axis direction. Therefore, the drag force F applied to the unfolded film 14 from the atmosphere in the X-axis direction has a component perpendicular to the roll axis (F in Figure 5). S It has (indicated by the sign of ). This results in 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 Center of gravity G S Join in around it. T S =d × F S =d × Fsinω This Torque T S This acts in the opposite direction to the tilt of spacecraft 2 around its yaw axis (Z axis), correcting the attitude of spacecraft 2 so that its roll axis is parallel to the X axis. In other words, torque T S This acts as a posture-restoring torque.

[0070] In contrast, the center of gravity G S ,G W Considering the example spacecraft 2' where no offset is set between them (see Figure 6), even if spacecraft 2 is tilted around the yaw axis (Z axis), the geometric centroid G of the deployed membrane 14 W The position is the center of gravity G of spacecraft 2. S Since this coincides with the position, the drag force F applied to the deployed film 14 does not have a component that rotates the spacecraft 2, and therefore does not act to correct the tilt of the spacecraft 2.

[0071] Unlike the spacecraft 2 shown in Figures 1, 3(A), 3(B), and 5, the geometric centroid G of the deployed membrane 14 is W The center of gravity of spacecraft 2 S If it is offset forward with respect to the direction of travel (X-axis), then, contrary to Figures 3(B) and 5, the center of gravity G of the spacecraft 2 will be at the deployed membrane 14 which is relatively forward with respect to the direction of travel (X-axis). S The torque generated in the deployable membrane 14, which is located relatively behind the center of gravity, is greater than the torque generated around the center of gravity, and the sum of the torques acts in a direction that increases the rotation of the spacecraft 2 (illustration omitted). Therefore, when offsetting the center of gravity of the spacecraft and the geometric center of gravity of the deployable membrane in order to obtain attitude restoration torque using a mechanism such as those shown in Figure 3(B) and Figure 5, it is necessary to position the geometric center of gravity of the deployable membrane so that it is located behind the center of gravity of the spacecraft with respect to the direction of travel (X-axis).

[0072] Furthermore, 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 geometric center of gravity of the deployable membrane is offset from the center of gravity of the spacecraft with respect to the roll axis, as described above, it is possible that at one point during flight, the geometric center of gravity of the deployable membrane may be located behind the center of gravity of the spacecraft with respect to the direction of travel (X-axis), but at another point, the geometric center of gravity of the deployable membrane may be located in front of the center of gravity of the spacecraft with respect to the direction of travel (X-axis).

[0073] Furthermore, the attitude restoration torque described above is equal to the geometric center of gravity G of the deployed film 14. W and the center of gravity G of spacecraft 2 SThe greater the offset between the centers of gravity (d in Figure 5), the greater the moment of inertia around the pitch axis and the yaw axis. If the moment of inertia around the pitch axis and the yaw axis is greater 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 will be greater. Therefore, it is preferable that the spacecraft 2 is designed such that, in the state where the deployable membrane 14 is deployed, the moment of inertia around the roll axis is greater than the moment of inertia around the other two axes (pitch axis and yaw axis) that intersect the roll axis.

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

[0075] The above-described settings for the moment of inertia are effective when the atmosphere is sufficiently dense and the effect of the attitude-restoring torque obtained by atmospheric drag is sufficiently large compared to the effect of gravity (for example, below medium altitude; generally less than 600 km). However, at high altitudes, such as above 600 km, the atmosphere is thin, so the drag force that spacecraft 2 experiences from the atmosphere decreases relatively in proportion to its speed. Therefore, depending on the shape and weight of spacecraft 2, the mechanical effect added by gravity in proportion to the moment of inertia becomes relatively larger compared to the effect of the attitude-restoring torque described above.

[0076] Under such conditions, depending on the relative magnitudes of the moment of inertia around each axis, spacecraft 2 may exhibit undesirable behavior due to the effects of gravity.

[0077] Specifically, if spacecraft 2 has a shape in which its moment of inertia around the pitch axis is smaller than its moment of inertia around the roll axis, and insufficient attitude-restoring torque is obtained due to atmospheric drag as described above, it is conceivable that spacecraft 2 may rotate around the yaw axis due to the effect of gravity. As a result, if the pitch axis of spacecraft 2 swings significantly towards the direction of travel, there is a concern that the attitude-restoring torque obtained from atmospheric drag will decrease even further, making it increasingly difficult to maintain attitude.

[0078] Therefore, research by the inventors of this invention has revealed that under conditions such as high altitude where atmospheric drag is relatively small, it is preferable that the moment of inertia around the pitch axis be sufficiently large. More specifically, it is preferable that spacecraft 2 be configured such that the moment of inertia around the pitch axis is greater than the moment of inertia around the roll axis.

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

[0080] To summarize, regarding the inertia of spacecraft 2, firstly, the inertia around the roll axis is greater than the inertia around the yaw axis (the inertia around the roll axis is greater than the yaw axis). xx , the moment of inertia around the yaw axis is I zz Therefore, I zz xx It can be said that 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.e., the moment of inertia around the pitch axis is greater than the moment of inertia around the yaw axis). yy Therefore, I zz xx And I yy xx ​​​It is preferable that (I zz xx yy It can be said that this is preferable.

[0081] The moment of inertia around each axis and the relative magnitudes of them can be adjusted, for example, by the offset amount of the geometric center of gravity of the unfolding membrane 14 with respect to the roll axis direction relative to the center of gravity of the main body 10, the angle of the unfolding membrane 14 with respect to each axis of the main body 10, and the dimensions of the unfolding membrane 14 that extend around each axis.

[0082] (Second Embodiment) In the first embodiment, the unfolded film 14 is unfolded along a plane perpendicular to the roll axis, and the geometric centroid G of the unfolded film 14 W The center of gravity of spacecraft 2 S The first embodiment describes a spacecraft 2 that is offset in relation to the roll axis direction. The second embodiment describes a spacecraft 2 in which the deployable membrane 14 is configured to be deployable to form an inclined surface that is sharply inclined with respect to the roll axis.

[0083] Figure 7 shows a spacecraft 2 according to the second embodiment. In the second embodiment of the spacecraft 2 shown in Figure 7, similar to the first embodiment shown in Figure 1, the deployment of the deployment membranes 14 from the movement control device 12 is configured such 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 direction of the roll axis. However, in this second embodiment, the movement control device 12 is equipped with an angle adjustment mechanism to change the angle of the deployment membranes 14, which is different from the first embodiment.

[0084] ​​As shown in Figure 7, the movement control device 12 is equipped with a mechanism (angle adjustment mechanism) that raises and lowers the fixing members 18 constituting each deployable membrane 14 relative to the main body 10 at the base end of the deployable membrane 14, allowing the angle of the deployable membrane 14 with respect to the roll axis to be adjusted. As a result, the spacecraft 2 can deploy the deployable membrane 14 along a plane perpendicular to the roll axis, as in the first embodiment shown in Figure 1, for example, but can also deploy it in a posture inclined at an acute angle with respect to the roll axis, as shown in Figure 7.

[0085] In the second embodiment of the spacecraft 2, the mechanism by which attitude restoration torque is generated in the inclined deployable membrane 14 will be explained with reference to Figure 8. Figure 8 is a schematic diagram showing the atmospheric drag force applied to the deployable membrane 14 when the roll axis of the spacecraft 2 is inclined with respect to the direction of travel (X axis) while the spacecraft 2 of the second embodiment is flying with the two deployable membranes 14 deployed so that the line segment connecting their centers of gravity is aligned with the yaw axis (Z axis).

[0086] For simplicity, this explanation will describe an example in which the two deployable membranes 14 are deployed by the movement control device 12 so as to form a plane passing through the center of gravity of the spacecraft 2, and are deployed rearward with respect to the roll axis (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 with respect to the center of gravity of the spacecraft 2, as in the first embodiment).

[0087] Let θ (0°C < θ < 90°) be the angle that the deployable membrane 14 makes with respect to the roll axis. Two deployable membranes 14 are deployed symmetrically with respect to the roll axis at an angle θ, and the spacecraft 2 is flying with its roll axis parallel to the X axis. As shown in Figure 8, the spacecraft 2 tilts at a slight angle φ around the pitch axis (Y axis), and assumes that one deployable membrane 14 (left side in the figure) is positioned further forward than the other (right side in the figure). In this case, the drag force that one deployable membrane 14 receives from the atmosphere perpendicular to the membrane surface 16 is F. L The cos(90°-θ+φ) force is cos(90°-θ+φ), and the drag force that the other developing film 14 receives from the atmosphere perpendicular to the film surface 16 is F. RThe torque is cos(90°-θ-φ). Therefore, if we let D be the distance between the center of gravity of spacecraft 2 and the center of gravity of each deployable membrane 14, then the torque T generated around the center of gravity of spacecraft 2 with respect to each deployable membrane 14 is cos(90°-θ-φ). Lt ,T Rt These are, T Lt =D×F Lt =D×F L cos(90°-θ+φ) T Lt =D×F Rt =D×F R cos(90°-θ-φ) This is the result.

[0088] If the left and right deployed membranes 14 in the figure are symmetrical with respect to the roll axis, the ratio of the projected areas of the left and right deployed membranes 14 in a plane perpendicular to the direction of collision of atmospheric molecules (X axis) is sin(θ-φ):sin(θ+φ). The drag force F exerted by the atmosphere on the left and right deployed membranes 14 is... L ,F R Since this is proportional to the ratio of the projected areas, 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 This acts in a direction that attempts to align the roll axis of spacecraft 2 with the X-axis, i.e., as an attitude-restoring torque.

[0089] In the second embodiment of the spacecraft 2, as shown in Figure 7, a movement control device 12 is provided that is configured to control the inclination angle of the deployable membrane 14. By adjusting the inclination angle of each deployable membrane 14, it is possible to control the magnitude of the attitude restoration torque as described above. For example, the following mechanism can be used to control the attitude restoration torque.

[0090] Figure 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 deployable membrane 14. The rotation mechanism 24 is provided with a rotation axis 44 oriented perpendicular to the roll axis, and the base end of the deployable membrane 14 is attached to the rotation axis 44 so that the deployable membrane 14 can be raised and lowered around the rotation axis 44. The movement control device 12 also includes a sensor 40, which is a torque sensor that detects the torque acting on the deployable membrane 14.

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

[0092] For example, the sensor 40 is mounted in the housing section 20 that houses the deployable film 14. When atmospheric molecules 34 collide with the deployable film 14 and atmospheric resistance acts upon it, a torque is applied to the deployable film 14, and the sensor 40 detects the torque transmitted from the deployable film 14 through the housing section 20. However, there are no restrictions on the mounting location of the sensor 40, as long as it can detect the torque acting on the deployable film 14. For example, the sensor 40 may be mounted directly on the deployable film 14 and housed together with the deployable film 14 in the housing section 20.

[0093] The rotation mechanism 24 rotates the deployable film 14 in accordance with the atmospheric resistance acting on the deployable film 14. The rotation mechanism 24 can arbitrarily set the angle of the deployable film 14 with respect to the rotation axis 44. Specifically, the rotation mechanism 24 is equipped with a rotation motor 42, and the deployable film 14, the housing section 20, and the sensor 40 are connected to the output shaft of the rotation motor 42. As the rotation motor 42, a servo motor, stepping motor, etc., that can arbitrarily set the rotation angle of the output shaft can be used. When the rotation motor 42 is operated, the deployable film 14, the housing section 20, and the sensor 40 rotate around the rotation axis 44.

[0094] Furthermore, the movement control device 12 includes a computer 46. The computer 46 corresponds to a control unit (control unit, control device) that controls the movement control device 12. Specifically, the computer 46 is connected to the components that make up the movement control device 12 (deployable film 14, sensor 40, 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 computer 46 is composed of a computer. In this case, the electronic computer 46 includes a processor such as a CPU (Central Processing Unit) that performs calculations and other processing necessary for controlling the movement control device 12, and a memory such as a ROM (Read Only Memory) or RAM (Random Access Memory) that stores various information (data, programs, etc.) used for controlling the movement control device 12. The electronic computer 46 controls the deployment and retraction of the unfolding film 14, the detection of torque by the sensor 40, and the rotation of the unfolding film 14 by the rotary motor 42 by executing the program stored in the memory.

[0096] The information detected by the sensor 40 is input to the computer 46. The computer 46 controls the rotary motor 42 based on the information detected by the sensor 40 to set the rotation angle of the unfolding film 14. In this way, the rotation angle of the unfolding film 14 is actively adjusted.

[0097] Alternatively, the control of the mobile control device 12 can be controlled using the computer mounted on the main body 10 of the spacecraft 2 (see Figure 1) instead of the computer 46. In this case, the computer 46 is omitted, and each component of the mobile control device 12 is connected to the computer provided on the main body 10. Control signals are then output from the computer on the main body 10 to each component of the main body 10 and the mobile control device 12, thereby controlling the operation of the main body 10 and the mobile control device 12.

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

[0099] The mobile control device 12 may also be powered by another battery located outside of it. For example, a battery located in the main body 10 of the spacecraft 2 (see Figure 1), or a battery located 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 of the spacecraft 2 (see Figure 1) even if the power from the battery located outside of the mobile control device 12 runs out or if it is difficult to supply power to the mobile control device 12 from an external source. Furthermore, the mobile control device 12 may receive power from both the battery 48 and a battery located outside of it.

[0100] Furthermore, the mobile control device 12 includes a generator 50 that generates power to operate the components of the mobile control device 12. For example, the 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 restrictions on the type of generator 50.

[0101] The generator 50 is appropriately installed in a position suitable for power generation by the movement control device 12. For example, the generator 50 may be provided in the housing 20, on the sensor 40, or on the rotating mechanism 24. In the example shown here, the generator 50 is shown 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 housing 20, the sensor 40, and the rotating motor 42, one or more generators 50 may be installed on the outer circumferential surface of the housing. In addition, the generator 50 can also be attached to the deployable membrane 14.

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

[0103] Furthermore, the mobile control device 12 may be equipped with a communication unit (not shown) for information communication. The communication unit is a communication module for sending and receiving information with any external communication device, and includes an antenna, a communication circuit, etc. By being equipped with a communication unit, the mobile control device 12 can communicate with the main body 10 of the spacecraft 2, other spacecraft, ground-based communication devices, etc.

[0104] As spacecraft 2 moves 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 sequentially detected by sensor 40, and the detected values ​​are input into computer 46. In this way, the force acting on the deployable membrane 14 is monitored.

[0105] Furthermore, the computer 46 outputs a control signal to the rotary motor 42 based on the force detected by the sensor 40, thereby controlling the rotary motor 42. The rotary motor 42 then sets the rotation angle of the unfolding film 14 to an angle corresponding to the force detected by the sensor 40, based on the control signal input from the computer 46.

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

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

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

[0109] By adjusting the rotation angle of the deployable membrane 14, the angle of the deployable membrane 14 with respect to the direction of travel of the spacecraft 2 can be adjusted, thereby adjusting the atmospheric drag acting on the deployable membrane 14. This allows for appropriate operations such as correcting the attitude of the spacecraft 2 or moving the spacecraft 2 out of orbit 6 (see Figure 1).

[0110] As described above, the movement control device 12 according to this second embodiment includes a rotation mechanism 24 that can arbitrarily set the angle of the membrane surface 16 of the deployable membrane 14 with respect to the roll axis. This makes it possible to adjust the rotation angle of the deployable membrane 14 so that a predetermined atmospheric resistance acts on the deployable membrane 14 even when the attitude of the spacecraft 2 changes, thereby enabling appropriate control of the attitude and movement of the spacecraft 2.

[0111] Furthermore, in the spacecraft 2 shown in Figures 7 and 8, if the deployable membrane 14 is deployed so as to be forward of the spacecraft 2 with respect to the roll axis (when θ > 90° in Figures 7 and 8), then, contrary to Figure 8, the torque generated around the center of gravity of the spacecraft 2 by the deployable membrane 14 located relatively behind the spacecraft 2 will be greater than the torque generated by the deployable membrane 14 located relatively behind the spacecraft 2 with respect to the X axis, and the sum of the torques will act in a direction that strengthens the rotation of the spacecraft 2 (illustration omitted). Therefore, when deploying the deployable membrane 14 diagonally with respect to the roll axis in order to obtain attitude restoration torque using the mechanism shown in Figure 7, the angle θ of the deployable membrane 14 needs to be acute with respect to the rear of the roll axis.

[0112] Furthermore, 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 deployable membrane has an inclined surface with respect to the spacecraft's roll axis, as described above, during flight, at one point the surface formed by the deployable membrane may be inclined at an acute angle with respect to the direction of travel (X-axis) relative to the rear of the roll axis, but at another point the surface formed by the deployable membrane may be inclined at an acute angle with respect to the direction of travel (X-axis) relative to the front of the roll axis (and obtuse angle with respect to the rear of the roll axis).

[0113] Furthermore, the attitude restoration torque described above increases as the angle θ that the deployable membrane 14 makes with respect to the roll axis decreases. On the other hand, as the angle θ decreases, the deployable membrane 14 protrudes further backward, and the diameter of the deployable membrane 14 around the roll axis decreases. As a result, the moment of inertia around the pitch axis and yaw axis increases, while the moment of inertia around the roll axis decreases. If the moment of inertia around the pitch axis and yaw axis is greater 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 becomes greater. Therefore, when deploying the deployable membrane 14, it is preferable to design the spacecraft 2 or adjust the angle of the deployable membrane 14 so that the moment of inertia around the roll axis is greater than the moment of inertia around the other two axes (pitch axis and yaw axis) that intersect the roll axis.

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

[0115] In other words, regarding the moment of inertia of spacecraft 2, first of all, the moment of inertia around the roll axis is greater than the moment of inertia around the yaw axis (I zz xx It is preferable that (I zz xx And I yy 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 xx yy It is preferable that it is.

[0116] ​​​​​In the case of the spacecraft 2 of this second embodiment, the angle of the deployable membrane 14 can be adjusted by the movement control device 12. This allows the offset amount of the geometric center of gravity of the deployable membrane 14 relative to the center of gravity of the main body 10 with respect to the roll axis, and the amount of extension of the deployable membrane 14 around the roll axis and pitch axis to be manipulated, thereby adjusting the moment of inertia around these axes. Using this mechanism, for example, in regions with relatively high altitudes, xx yy Therefore, in regions with relatively low altitudes, I yy xx Theoretically, it is also possible to operate the movement control device 12 in such a way.

[0117] The relationship between the inclination of the surface formed by the deployed film 14 and the attitude restoration torque is as described above. However, in order to generate the attitude restoration torque as described above, the deployed film 14 does not necessarily need to have an entire film surface 16 that forms an inclined surface extending backward at an acute angle with respect to the roll axis; such an inclination may be provided only in part. For example, part of the film surface 16 may be a surface perpendicular to the roll axis, while the rest has an inclined surface as described above. Furthermore, the film surface 16 may be a curved surface, or it may have surfaces with different inclination angles in different parts.

[0118] Here, we will also explain the case where the spacecraft 2 of the second embodiment is tilted in a different direction with respect to the X-axis (around the yaw axis) with reference to Figure 10.

[0119] Figure 10 shows the second embodiment of the spacecraft 2 in flight with its two deployable membranes 14 deployed such that the line segment connecting their centers of gravity aligns with the yaw axis (Z axis), and the spacecraft is tilted at a slight angle ω with respect to the X axis around the yaw axis. In this case, if the cross-section of the surface formed by the deployable membranes 14 along the XY plane (roll axis-pitch axis plane) is perpendicular to the roll axis when viewed from the yaw axis (Z axis), then the drag force F that the deployable membranes 14 receive from the atmosphere along the X axis acts parallel to the roll axis on the deployable membranes 14.

[0120] ​​Therefore, as shown in Figure 8, for example, the direction of the resistive force applied to the deployable membrane 14 is not different on the left and right sides of the figure. However, if the center of gravity of the deployable membrane 14 is offset aft 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 almost the same as that shown in Figure 5 (the component of the resistive force acting as this attitude restoring torque is shown in Figure 10 as F S (Indicated by the symbol ).

[0121] In manufacturing the spacecraft 2 on which the attitude restoration torque described above acts, a spacecraft body 10 and a movement control device 12 having a deployable membrane 14 are prepared, and the movement control device 12 is attached to the body 10 such that the geometric center of gravity of the deployed membrane 14 is offset from the center of gravity of the body 10 with respect to the roll axis. At that time, the movement control device 12 is attached to the body 10 such that the deployed membrane 14 forms an acute angle with respect to the rear of the roll axis.

[0122] Furthermore, the spacecraft 2 of the first embodiment as described in Figures 3(B) and 5 (center of gravity G of spacecraft 2) S In contrast, the geometric centroid G of the unfolded film 14 W The attitude restoration torque in the case where the position is offset rearward with respect to the X-axis direction, and the attitude restoration torque in the second embodiment of the spacecraft 2 as described in Figure 8 (where the surface formed by the deployable membrane 14 extends rearward with respect to the roll axis from the base end to the tip and is inclined at an acute angle with respect to the roll axis), have different generation principles and can occur independently of each other.

[0123] Therefore, for example, in a spacecraft where the geometric center of gravity of the deployable membrane is located forward of the spacecraft's center of gravity in the direction of the roll axis (configuration A), and the surface formed by the deployable membrane extends backward from the base end towards the tip in the direction of the roll axis and is inclined at an acute angle with respect to the roll axis (configuration B), if the attitude restoration torque from configuration B exceeds the torque that tends to increase the tilt of the spacecraft caused by configuration A when the roll axis of the spacecraft tilts with respect to the X axis during flight, then the attitude of the spacecraft can be corrected.

[0124] Conversely, in a spacecraft where the geometric center of gravity of the deployable membrane is located aft of the spacecraft's center of gravity in the direction of the roll axis (referred to as configuration C), and the surface formed by the deployable membrane extends forward from the base end to the tip in the direction of the roll axis while forming an inclined surface with respect to the roll axis (referred to as configuration D), if the attitude restoration torque from configuration C exceeds the torque generated by configuration D that tends to increase the tilt of the spacecraft when the roll axis of the spacecraft tilts with respect to the X axis during flight, then the attitude of the spacecraft can be corrected.

[0125] (Third embodiment) Figure 11 is a perspective view showing the configuration of the spacecraft 2 according to the third embodiment. In the spacecraft 2 of the third embodiment, similar to the second embodiment, 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 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 this third embodiment, each deployable membrane 14 can be rotated about the axis of the deployable membrane 14 (an axis that follows the plane formed by the deployable membrane 14 and is 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 the same as that of the rotation mechanism 24 in the second embodiment described above (see Figure 9), but the orientation of the rotation axis 44 is different. That is, in the spacecraft 2 of this third embodiment, the orientation of the rotation axis 44 coincides with the axis of the deployable membrane 14, and the deployable membrane 14 rotates around the axis which is 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 for detecting torque acting on the deployable membrane 14, a rotary motor 42 for rotating the deployable membrane 14, a computer 46 for controlling the movement control device 12, a battery 48 for supplying power to the components of the movement control device 12, a generator 50, a communication unit, etc. (see Figure 9). The torque acting on the deployable membrane 14 due to atmospheric resistance is detected by the sensor 40, and the angle of the deployable membrane 14 can be adjusted based on this.

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

[0129] The central axis of the cylindrical body 10 extends longer than its diameter. That is, the longitudinal direction of the body 10 coincides with the axial direction. Furthermore, the body 10 has a mass that is sufficiently large compared to the mass of the unfolded film 14.

[0130] In this type of spacecraft 2, due to the action of gravity, the longitudinal direction of the main body 10 tends to align with the direction of gravity. That is, in the spacecraft 2 of this fourth embodiment, the direction along the longitudinal direction of the columnar main body 10 becomes the yaw axis. Furthermore, regarding the moment of inertia, the moment of inertia around the yaw axis (I zz ) has a moment of inertia (I) around the other two axes (roll axis and pitch axis). xx , I yy It becomes smaller than ). Because the main body 10, which has a large mass, extends along the yaw axis, xx and I yy This is because it has become sufficiently large. In other words, a lightweight member such as the unfolding membrane 14, compared to the main body 10, extends around the yaw axis, and this I zz Even if it increases somewhat, zz I xx and I yy It does not surpass that.

[0131] The spacecraft 2 of the fourth embodiment comprises two deployable membranes 14. Each deployable membrane 14 is attached at its base end to a rearward position on the side of the cylindrical spacecraft 2, as shown in Figures 12 and 13, and extends diagonally backward from there.

[0132] In a front view (a field of view of the spacecraft 2 from the front along the roll axis), the two deployable membranes 14 are positioned to extend 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 and yaw axis direction), and extend backward from the main body 10 at oblique angles with respect to the XZ plane (vertical plane including the roll axis and yaw axis) and the YZ plane (vertical plane including the pitch axis and yaw axis). Thus, each membrane surface 16 forms a plane that intersects obliquely with respect to the direction of travel of the spacecraft 2 (direction along the roll axis), and is subject to atmospheric resistance associated with the flight of the spacecraft 2.

[0133] Geometric centroid G of the two unfolded films 14 w It is located in a position offset to the rear 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 of the two unfolding membranes 14 w This is the center of gravity G of the entire spacecraft 2. s It is also located aft with respect to the roll axis.

[0134] The configuration of the unfolding film 14 shown herein is merely an example, and the number and shape of the unfolding film 14, as well as its mounting position and angle relative to the main body 10, can be set as appropriate in addition to the example shown herein.

[0135] Regarding the setting of the moment of inertia, as in the first and second embodiments described above, I zz xx It is preferable that this be the case, but in this third embodiment in particular, due to the shape and weight of the main body 10, it is inevitable that I zz xx This is the case. Under conditions where the altitude is relatively low and the atmosphere is relatively dense, I zz xx And I yy xx It is preferable that this is the case. Under conditions where the altitude is relatively high and the atmosphere is relatively thin, I​​​​zz xx yy It is preferable that this be the case.

[0136] The moment of inertia around each axis and the relative magnitudes of them can be adjusted, for example, by the offset amount of the geometric center of gravity of the unfolding membrane 14 with respect to the roll axis direction relative to the center of gravity of the main body 10, the angle of the unfolding membrane 14 with respect to each axis of the main body 10, and the dimensions of the unfolding membrane 14 that extend around each axis.

[0137] Furthermore, the movement control device 12 of the spacecraft 2 may be equipped with a rotation mechanism 24 similar to that in the second and third embodiments described above, so as to be able to change the angle of the deployed membrane 14 relative to the main body 10.

[0138] (Fifth embodiment) Figure 14 is a perspective view showing the form of the spacecraft 2 according to the fifth embodiment, and Figure 15 is a side view showing the form of the spacecraft 2 in Figure 14. The spacecraft 2 of this fifth embodiment, like the spacecraft 2 of the fourth embodiment (see Figures 12 and 13), comprises a cylindrical body 10 and a movement control device 12 having two deployable membranes 14.

[0139] The main body 10 has a mass sufficiently large compared to the mass of the unfolding film 14, and extends along the yaw axis with the central axis of its cylindrical shape oriented in the longitudinal direction. The base ends of the two unfolding films 14 are attached near the upper end of this main body 10.

[0140] In a front view (a field of view of the spacecraft 2 from the front along the roll axis), the two deployable membranes 14 are positioned to extend downward and 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 generally parallel to the pitch axis direction and extend backward and downward from the main body 10 at oblique angles with respect to the XY plane (vertical planes including the roll axis and pitch axis) and the YZ plane (vertical planes including the pitch axis and yaw axis). Thus, each membrane surface 16 forms a plane that intersects obliquely with respect to the direction of travel of the spacecraft 2 (direction along the roll axis) and is subject to atmospheric resistance associated with the flight of the spacecraft 2.

[0141] ​​Geometric centroid G of the two unfolded films 14 w It is located in a position offset to the rear 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 of the two unfolding membranes 14 w This is the center of gravity G of the entire spacecraft 2. s It is also located aft with respect to the roll axis.

[0142] In this way, the mounting position and angle of the deployable 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 deployed deployable membrane 14 has an inclined surface that forms an acute angle with respect to the roll axis. As a result, when the attitude of the spacecraft 2 changes, an attitude restoration torque is applied to correct the attitude change, making it possible to appropriately control the attitude and movement of the spacecraft 2.

[0144] The configurations, methods, etc., of each embodiment described above can be modified as appropriate without departing from the scope of the object of the present invention. Furthermore, the configurations, methods, etc., of this embodiment can be combined with other embodiments as appropriate. [Explanation of Symbols]

[0145] 2 Spacecraft 4 celestial bodies 6 orbits 8 Detachment Trajectory 10 Main Unit 12 Mobile control device 14. Developing film (resistive film) 16 Membrane surface 18 Fixing member 20 Storage Units 24 Rotation mechanism 34 Atmospheric molecules 40 sensors 42 Rotation motor 44 Rotation axis 46 Electronic computer 48 batteries 50 Generators

Claims

1. The spacecraft itself, 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 pair of fixing members to which the membrane surface is fixed are spaced further apart from each other as they move away from the spacecraft, and the membrane surface is formed to be wider as it moves away from the spacecraft body. The 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 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.

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

5. The spacecraft according to any one of claims 1 to 4, wherein the mobile control device further comprises one or both of a solar cell and a battery.

6. The spacecraft according to any one of claims 1 to 4, wherein the movement control device causes the spacecraft body to leave orbit by deploying the deployment membrane.

7. The spacecraft itself, A moving control device is prepared, which has a deployable membrane that is deployed so as to be subjected 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, The pair of fixing members to which the membrane surface is fixed are spaced further apart from each other as they move away from the spacecraft, and the membrane surface is formed to be wider as it moves away from the spacecraft body. 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.

8. The method for manufacturing a spacecraft according to claim 7, wherein the movement control device is attached to the spacecraft body such that the deployed membrane, when deployed, forms an acute angle with respect to the rear of the roll axis.

9. The spacecraft itself, 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 pair of fixing members to which the membrane surface is fixed are spaced further apart from each other as they move away from the spacecraft, and the membrane surface is formed to be wider as it moves away from the spacecraft body. 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.

10. The spacecraft itself, A moving control device is prepared, which has a deployable membrane that is deployed so as to be subjected 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, The pair of fixing members to which the membrane surface is fixed are spaced further apart from each other as they move away from the spacecraft, and the membrane surface is formed to be wider as it moves away from the spacecraft body. A method for manufacturing a spacecraft, comprising attaching the movement control device to the spacecraft body such that the deployed membrane has an inclined surface that forms an acute angle with respect to the roll axis when deployed.

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