A space non-cooperative target orbit transfer system and method based on tether momentum isolation

Abstract

The application discloses a space non-cooperative target orbit transfer system and method based on tether momentum isolation, and belongs to the technical field of tether satellite systems and space orbit transfer. The application solves the problems of the existing orbit transfer method, i.e. affecting the normal operation of a spacecraft and large energy consumption. The application adopts a space tether structure, releases a tether orbit transfer device, and collides with a non-cooperative target at a long distance to transfer the orbit of the non-cooperative target. The adverse momentum generated by the collision is isolated by a soft conductive tether and is gradually reduced during the recycling process of the orbit transfer device, so that the normal operation of the spacecraft in orbit is not affected. During the recycling process of the tether orbit transfer device, the Lorentz force of the conductive tether in the earth's magnetic field can offset part of the adverse angular momentum, reduce the dependence on a propeller, and reduce the energy consumption of the system. In addition, the current generated by the conductive tether cutting the magnetic induction lines can charge the orbit transfer device, so that energy recovery is realized. The method can be applied to the orbit transfer task of a space non-cooperative target.

Description

Technical Field

[0001] The present invention belongs to the technical field of tethered satellite systems and space orbit transfer, and in particular relates to a space non-cooperative target orbit transfer system and method based on tether momentum isolation. Background Art

[0002] The dramatic increase in the number of non-cooperative objects in space, such as space debris and space junk, has seriously jeopardized the normal operation of spacecraft in orbit, causing damage to insulation, critical components, functional limitations, or even failure. Tethered satellite systems, as a new type of spacecraft system, connect multiple spacecraft with tethers to enable large-scale space maneuvers. They are expected to become an effective solution for many space missions, especially for reorienting non-cooperative objects and preventing them from disrupting orbital spacecraft.

[0003] To counter the intrusion of non-cooperative space targets on in-orbit spacecraft, traditional solutions include passive and active solutions. The former requires installing protective materials on in-orbit spacecraft or using on-orbit maneuvers to evade intrusion; the latter requires irradiating non-cooperative targets with high-energy lasers, causing them to sublimate or shift their orbits. However, installing protective materials on spacecraft can only defend against smaller non-cooperative targets, and the defense process will generate momentum impact on the spacecraft, affecting its motion state. Secondly, while on-orbit maneuvers are an effective means of defending against non-cooperative targets, they will affect the normal operation of the spacecraft during the evasion process and waste a lot of energy. Finally, using high-energy lasers to eliminate non-cooperative targets also has the problem of generating momentum impact on the spacecraft and consuming a lot of energy.

[0004] Therefore, there is an urgent need for a space non-cooperative target orbit transfer system and method that is efficient, reliable, has active defense, has low operating costs, and does not affect the operating status of on-orbit spacecraft. Summary of the Invention

[0005] The purpose of the present invention is to solve the problem that the existing method will affect the normal operation of the spacecraft and consume a lot of energy, and to propose a space non-cooperative target orbit transfer system and method based on tether momentum isolation.

[0006] The technical solution adopted by the present invention to solve the above technical problems is:

[0007] According to one aspect of the present invention, a space non-cooperative target orbit transfer system based on tethered momentum isolation is provided, wherein the space non-cooperative target orbit transfer system comprises an on-orbit spacecraft and a tethered orbit transfer device, wherein the on-orbit spacecraft and the tethered orbit transfer device are connected via a conductive tether 8;

[0008] The tethered track transfer device includes an impact plate 1, an energy storage structure, a photoelectric sensor 4, a tethered track transfer device body 6, a Spindt cathode emitter 7 and a vector thruster 10;

[0009] After the conductive tether 8 passes through the interior of the tethered track transfer device body 6, it is connected to the Spindt cathode emitter 7 on the surface of the tethered track transfer device body 6;

[0010] The main body 6 of the rope-tethered track transfer device is provided with an impact plate 1, a photoelectric sensor 4 and a vector thruster 10, and an energy storage structure is provided at a symmetrical position on the opposite surface of the impact plate 1 and the main body 6 of the rope-tethered track transfer device;

[0011] The energy storage structure includes a sleeve 2, a spring 3 and an electromagnet;

[0012] The sleeve 2 is symmetrically arranged on the opposite surfaces of the rope track transfer device body 6 and the impact plate 1, and a hollow electromagnet is arranged at the non-fixed end of the sleeve 2;

[0013] After the spring 3 passes through the symmetrically arranged sleeve 2 and the electromagnet, one end of the spring 3 is fixedly connected to the rope track transfer device body 6, and the other end is fixedly connected to the impact plate 1.

[0014] Furthermore, four energy storage mechanisms are provided between the rope-tethered track transfer device body 6 and the impact plate 1 .

[0015] Furthermore, the impact plate 1 is a circular plate with a partially hollowed area, and a photoelectric sensor 4 is provided on the upper surface of the impact plate 1;

[0016] The photoelectric sensor 4 on the rope-tethered track transfer device body 6 and the impact plate 1 is used to observe and sense the relative position of the non-cooperative target and the rope-tethered track transfer device.

[0017] Furthermore, the rope-track transfer device further includes a first foldable solar wing 5 and a second foldable solar wing 9 , and the first foldable solar wing 5 and the second foldable solar wing 9 are hinged on the rope-track transfer device body 6 .

[0018] Furthermore, the on-orbit spacecraft is provided with a tether release module, a release speed adjustment module and a steering adjustment unit;

[0019] The conductive tether 8 is connected to the tether release module, which is used to release the tether track transfer device. The release speed adjustment module is used to adjust the release speed of the track transfer device. The steering adjustment unit is used to adjust the release angle of the track transfer device.

[0020] Furthermore, the rope-tethered track transfer device is provided with a vector control device, and the vector control device is used to provide the rope-tethered track transfer device with torque within the track plane and outside the track plane.

[0021] According to another aspect of the present invention, a method for space non-cooperative target orbit transfer based on tether momentum isolation specifically comprises the following steps:

[0022] Step S1: Select a non-cooperative space target that needs to perform an orbit transfer mission based on the relative orbit information between the on-orbit spacecraft and the non-cooperative space target; then plan and calculate the initial velocity and initial angle of the tethered orbit transfer device after release, and the orbital maneuver trajectory of the tethered orbit transfer device after release;

[0023] Step S2: Using the steering adjustment unit and release speed adjustment module on the on-orbit spacecraft to adjust the release of the tethered orbit transfer device according to the initial angle and initial speed. When the orbit transfer device is moving away from the on-orbit spacecraft, the tether release module on the on-orbit spacecraft controls the tension on the tether to adjust the release speed of the orbit transfer device. Using the vector control device, the system is provided with torque in and out of the orbital plane, thereby adjusting the system's in-plane and out-of-plane angles and controlling the orbit transfer device to operate according to the calculated orbital maneuver trajectory.

[0024] Step S3: Using the photoelectric sensor on the tethered track transfer device to detect the position of the non-cooperative target, and then using the vector thruster 10 on the track transfer device to adjust the track transfer device posture so that the impact plate 1 faces the non-cooperative target. Then, using the electromagnetic catapult collision device to send a release signal, after the spring is released, the impact plate 1 hits the non-cooperative target;

[0025] Step S4: Recover the conductive tether 8 and pull the tether orbit transfer device to recover it to the on-orbit spacecraft.

[0026] Furthermore, the non-cooperative space target that needs to perform the orbit transfer mission is selected based on the relative orbit information between the on-orbit spacecraft and the non-cooperative space target; the specific method adopts the following method (1) or method (2);

[0027] Method (1):

[0028] The positions of non-cooperative space targets in the Earth's inertial coordinate system are inferred based on their catalog information. Combined with the positions of on-orbit spacecraft in the Earth's inertial coordinate system, the relative orbital information between the on-orbit spacecraft and each non-cooperative space target is calculated.

[0029] Based on the relative orbital information between the on-orbit spacecraft and various space non-cooperative targets, the space non-cooperative targets that have a collision risk with the on-orbit spacecraft are determined. The determined space non-cooperative targets are the space non-cooperative targets that require orbit transfer missions.

[0030] Method (2):

[0031] The relative position information between each non-cooperative space target and the on-orbit spacecraft is measured using the sensing mechanism installed on the on-orbit spacecraft;

[0032] According to the relative orbital information between the on-orbit spacecraft and various space non-cooperative targets, the space non-cooperative targets that have a collision risk with the on-orbit spacecraft are determined. The determined space non-cooperative targets are the space non-cooperative targets that need to perform orbit transfer missions.

[0033] Furthermore, in step S2, the tether release module on the on-orbit spacecraft is used to control the tension on the tether to adjust the release speed of the tethered orbit transfer device. The vector control device is used to provide the system with torque in and out of the orbital plane to control the tethered orbit transfer device to operate according to the calculated orbital maneuver trajectory. The specific process is as follows:

[0034] Step S21: The model of the space tether system consisting of the on-orbit spacecraft, tether, and tether orbit transfer device is constructed as follows:

[0035]

[0036] Where q = (λ,θ,φ) T , λ is the normalized parameter of the system rope length, is the first derivative of q, is the second derivative of q, θ is the orbital plane angle of the system, φ is the orbital plane angle of the system, u=(u l ,u θ ,u φ ) T ,u l is the normalized control variable of the tether tension, u θ is the normalized control quantity of the torque in the orbital plane of the system, u φ is the normalized control quantity of the system's out-of-plane torque, is the first-order coefficient matrix of the system, G(q) is the zero-order coefficient matrix of the system, and H(q) is the controller input coefficient matrix;

[0037] G(q) and H(q) are:

[0038]

[0039] Where e is the eccentricity of the Keplerian orbit of the system, v is the true near angle of the system, k is the orbit coefficient, and ρ is the bias coefficient;

[0040] Step S22: Design the sliding surface structure according to the spatial tether system model:

[0041]

[0042] Among them, s i (q i ) is the sliding surface, q1=λ-1, q2=θ, q3=φ, It is q i The first derivative of |q i | isq i The absolute value of sgn(·) is the sign function, k i1 ,k i2 ,α i1 ,γ i1 is the sliding surface coefficient, and the sliding surface coefficient satisfies k i1 >0,k i2 >0,0<γ i1 <1<α i1 ;

[0043] The controller output is:

[0044]

[0045] in, is a matrix The i-th row element, G i (q) is the i-th row element of G(q), H(q) ii -1 refers to H(q) -1 The element in row i and column i in , H(q) -1 is the inverse matrix of H(q), k i3 ,k i4 ,α i2 is the controller coefficient, and satisfies k i3 >0,k i4 >0,α i2 >1;

[0046] Step S23 : The tethered release module and the vector control device control the tethered track transfer device to reach the location of the non-cooperative target along the track maneuvering trajectory according to the output of the controller.

[0047] Furthermore, the specific process of step S4 is as follows:

[0048] Step S41: using the electromagnetic thruster on the tethered track transfer device to eject high-speed charged ions to apply positive thrust to the tethered track transfer device;

[0049] Step S42: When the tethered track transfer device pulled by the conductive tether 8 is in the recovery process, the conductive tether 8 will cut the Earth's magnetic field, that is, electrons will accumulate on the conductive tether 8. The accumulated electrons are then emitted by the Spindt cathode emitter installed on the tethered track transfer device to form an electric current;

[0050] The current in step S43 and step S42 moves in the Earth's magnetic field to generate a Lorentz force, and the tethered track transfer device is braked using the Lorentz force and the forward thrust in step S41;

[0051] Step S44: Utilize the tether release module on the on-orbit spacecraft to recover the conductive tether 8, and pull the tether orbit transfer device to recover it to the on-orbit spacecraft.

[0052] The beneficial effects of the present invention are:

[0053] This paper proposes a method for orbit transfer of non-cooperative targets in space based on tether momentum isolation, which can achieve long-distance orbit transfer of non-cooperative targets and ensure the stable and efficient operation of on-orbit spacecraft:

[0054] 1. Long-range collision and deflection of non-cooperative targets in space, with a long safety distance. This invention utilizes a space tether structure, releasing the tethered trajectory shifting device to perform a long-range kinetic collision on non-cooperative targets, shifting their trajectory. The tether's release distance can reach several thousand to tens of kilometers, enabling the interception and deflection of non-cooperative targets at a considerable distance, maintaining a considerable safety distance.

[0055] 2. The orbital transfer operation is momentum-isolated from the on-orbit spacecraft, preventing it from operating normally. Because the two are connected by a flexible conductive tether, the momentum generated by the tethered orbital transfer device's collision with a non-cooperative object in space is not transferred to the on-orbit spacecraft. Instead, it is isolated by the flexible conductive tether and gradually dissipated during the tethered orbital transfer device's recovery process, resulting in no impact or interference with the normal operation of the on-orbit spacecraft.

[0056] 3. Using kinetic collisions to achieve orbital transfer of non-cooperative space targets consumes minimal energy. The collision mechanism on the tethered orbit transfer device alters the trajectory of non-cooperative space targets, forcing them to change speed, change trajectory, or even re-enter the atmosphere for destruction. This consumes minimal energy compared to traditional spacecraft maneuvers and laser removal methods. Furthermore, the drag in the space environment is very low, so orbital transfer operations with the tethered orbit transfer device do not require additional energy.

[0057] 4. The conductive tether can assist in stabilizing the tethered orbital transfer device and enable energy recovery. This invention utilizes a conductive tether structure. During the recovery process of the tethered orbital transfer device, the Lorentz force exerted by the conductive tether in the Earth's magnetic field can offset some of the unfavorable angular momentum, reducing the recovery process's reliance on the thruster and lowering system energy consumption. Furthermore, the induced current generated by the conductive tether cutting through the magnetic flux lines can charge the tethered orbital transfer device, enabling energy recovery and reuse. BRIEF DESCRIPTION OF THE DRAWINGS

[0058] Figure 1 This is a schematic structural diagram of a space non-cooperative target orbit transfer system based on tether momentum isolation according to the present invention;

[0059] Figure 2 It is a microscopic diagram of the Spindt cathode emitter 7;

[0060] Figure 3 1 is a schematic structural diagram of the impact plate 1;

[0061] Figure 4 It is a schematic diagram of the ejection collision process;

[0062] Figure 5 It is a schematic diagram of the orbit transfer process of a non-cooperative target in space;

[0063] Figure 6 It is a flow chart of the method of the present invention. DETAILED DESCRIPTION

[0064] Specific implementation method 1: Combination Figure 1 This embodiment is described. This embodiment describes a space non-cooperative target orbit transfer system based on tethered momentum isolation, comprising an on-orbit spacecraft and a tethered orbit transfer device, wherein the on-orbit spacecraft and the tethered orbit transfer device are connected via a conductive tether 8.

[0065] The tethered track transfer device includes an impact plate 1, an energy storage structure, a photoelectric sensor 4, a tethered track transfer device body 6, a Spindt cathode emitter 7 and a vector thruster 10;

[0066] One end of the conductive tether 8 connected to the tethered track transfer device is located inside the tethered track transfer device body 6; and the conductive tether 8 inside the tethered track transfer device body 6 is connected to the Spindt cathode emitter 7 on the surface of the tethered track transfer device body 6;

[0067] The impact plate 1 is arranged above the main body 6 of the tethered track transfer device, the vector thruster 10 is arranged at the bottom of the main body 6 of the tethered track transfer device, and the photoelectric sensor 4 is installed around the main body 6 of the tethered track transfer device;

[0068] The energy storage structure includes a sleeve 2, a spring 3 and an electromagnet;

[0069] Sleeves 2 are symmetrically arranged on the upper surface of the rope-tethered track transfer device body 6 and the lower surface of the impact plate 1, and hollow electromagnets are arranged at the non-fixed ends of the sleeves 2;

[0070] After the spring 3 passes through the symmetrically arranged sleeve 2 and electromagnet, one end of the spring 3 is fixedly connected to the upper surface of the rope track transfer device body 6, and the other end is fixedly connected to the lower surface of the impact plate 1.

[0071] The lower end of the sleeve 2 provided on the lower surface of the impact plate 1 is connected to an electromagnet, and the upper end of the sleeve 2 provided on the upper surface of the rope track transfer device body 6 is also connected to an electromagnet. In this way, the sleeve can be electrically attracted to compress the spring to store energy, and the sleeve can be discharged to release the energy. The spring triggers the impact plate 1 to pop out and hit the non-cooperative target. Figure 4 As shown, in the initial stage, the orbit transfer device is in state one, the spring is in a compressed state and is located in a sleeve on the orbit transfer device; when a non-cooperative target in space is detected, the electromagnet is powered off, and the spring is released by relying on elastic potential energy, and the spring is in a state of continuous rebound, i.e., state two; in the collision stage, the orbit transfer device is in state three, and the impact plate 1 collides with the non-cooperative target in space, performs the collision function, completes the potential energy-kinetic energy conversion, and realizes the orbit transfer of the non-cooperative target. In the recovery stage, the orbit transfer device is in state four, and the impact plate 1 rebounds under the action of the reaction force of the non-cooperative target in space and the potential energy of the spring, compressing the spring. When the impact plate moves to a position where the electromagnet can be attracted, the electromagnet is instantly powered on, and the system returns to state one. The method of the present invention is used to transfer the orbit of a non-cooperative target in space, which can isolate the momentum of the on-orbit spacecraft and prevent the working state of the on-orbit spacecraft from being affected.

[0072] The Spindt cathode emitter 7 used in the present invention is an existing structure, such as Figure 2 As shown, the Spindt cathode emitter 7 consists of a thin upper metal gate layer and a lower silicon base sandwiched by a silicon dioxide insulating layer, allowing electrons to be freely emitted. An array of metal tips is distributed across the upper conductive and insulating layers. This emitter is compact, and the gate voltage between the upper and lower layers precisely controls the electron emission of each tip. Each tip is capable of emitting 100μA of current.

[0073] Specific embodiment 2: This embodiment differs from specific embodiment 1 in that four energy storage mechanisms are provided between the rope-tethered track transfer device body 6 and the impact plate 1 .

[0074] Other steps and parameters are the same as those in the first embodiment.

[0075] Specific implementation method three: Combination Figure 3 This embodiment is different from the first or second embodiment in that the impact plate 1 is a circular plate with a partially hollowed area, and a photoelectric sensor 4 is provided on the upper surface of the impact plate 1;

[0076] The photoelectric sensor 4 on the rope-tethered track transfer device body 6 and the impact plate 1 is used to observe and sense the relative position of the non-cooperative target and the rope-tethered track transfer device.

[0077] Other steps and parameters are the same as those in the first or second embodiment.

[0078] The hollow structure is mainly used to filter out tiny dust particles that consume the flight momentum of the rope-tethered track transfer device.

[0079] Specific embodiment four: This embodiment differs from any one of specific embodiments one to three in that the rope-tethered track transfer device further includes a first foldable solar wing 5 and a second foldable solar wing 9, and the first foldable solar wing 5 and the second foldable solar wing 9 are hinged on the rope-tethered track transfer device body 6.

[0080] The other steps and parameters are the same as those in the first to third embodiments.

[0081] The first foldable solar wing 5 and the second foldable solar wing 9 can be folded during the release flight of the tethered track transfer device, and unfolded to absorb energy during the standby waiting process.

[0082] Specific embodiment 5: This embodiment differs from any one of specific embodiments 1 to 4 in that the on-orbit spacecraft is provided with a tether release module, a release speed adjustment module and a steering adjustment unit;

[0083] The conductive tether 8 is connected to the tether release module, which is used to release the tether track transfer device. The release speed adjustment module is used to adjust the release speed of the track transfer device. The steering adjustment unit is used to adjust the release angle of the track transfer device.

[0084] The other steps and parameters are the same as those in the first to fourth embodiments.

[0085] Specific embodiment six: This embodiment differs from any one of specific embodiments one to five in that the rope-tethered track transfer device is provided with a vector control device, and the vector control device is used to provide the rope-tethered track transfer device with torque within and outside the track plane.

[0086] The other steps and parameters are the same as those in the first to fifth embodiments.

[0087] Specific implementation method seven: combination Figure 5 and Figure 6 This embodiment describes a method for space non-cooperative target orbit transfer based on tether momentum isolation, which specifically includes the following steps:

[0088] Step S1: Select a non-cooperative space target that needs to perform an orbit transfer mission based on the relative orbit information between the on-orbit spacecraft and the non-cooperative space target; then plan and calculate the initial velocity and initial angle of the tethered orbit transfer device after release, and the orbital maneuver trajectory of the tethered orbit transfer device after release;

[0089] Step S2: Using the steering adjustment unit and release speed adjustment module on the on-orbit spacecraft to adjust the release of the tethered orbit transfer device according to the initial angle and initial speed. When the orbit transfer device is moving away from the on-orbit spacecraft, the tether release module on the on-orbit spacecraft controls the tension on the tether to adjust the release speed of the orbit transfer device. Using the vector control device to provide forces in and out of the orbital plane to the system, thereby adjusting the system's in-plane and out-of-plane angles and controlling the orbit transfer device to operate according to the calculated orbital maneuver trajectory.

[0090] Step S3: Using the photoelectric sensor on the tethered track transfer device to detect the position of the non-cooperative target, then using the vector thruster 10 on the track transfer device to adjust the track transfer device so that the impact plate 1 faces the non-cooperative target, and then using the electromagnetic catapult collision device to send a release signal. After the spring is released, the impact plate 1 hits the non-cooperative target;

[0091] Step S4: Recover the conductive tether 8 and pull the tether orbit transfer device to recover it to the on-orbit spacecraft.

[0092] In step S3 of the present invention, after the tethered orbit transfer device arrives at the designated location, photoelectric sensors mounted around the orbit transfer device detect reflected light signals, using the reflected light to determine the location of the non-cooperative target in space. After determining the non-cooperative target's position, the orbit transfer device's vector thrusters 10 are used to adjust the orbit transfer device's position so that the impact plate 1 is facing the non-cooperative target. At this point, a release signal is sent to the electromagnetic catapult collision device. After the electromagnetic catapult device is unlocked, the impact plate 1 is ejected and strikes the non-cooperative target, causing the non-cooperative target to deviate from its original trajectory and no longer pose a threat to the spacecraft's operation. The electromagnetic catapult collision device consists of an electromagnet and four high-strength springs connected to a metal impact plate. When the electromagnet is energized, it generates suction to compress the springs, causing the impact plate 1 to retract. When the release signal is received, the electromagnet deactivates the electromagnetic force, dissipating the springs. The springs release, causing the impact plate 1 to eject at high speed and collide with the non-cooperative target, changing its trajectory. This causes the non-cooperative target to change speed and trajectory, or even re-enter the atmosphere and be destroyed. After the impact plate 1 completes the impact, it rebounds under the action of the reaction force of the non-cooperative target and the potential energy of the spring, and the electromagnet is instantly powered on, re-locked, and restored to the initial state.

[0093] Specific embodiment eight: This embodiment differs from specific embodiment seven in that the non-cooperative space target that needs to perform the orbit transfer mission is selected based on the relative orbit information between the on-orbit spacecraft and the non-cooperative space target; the specific method adopts the following method (1) or method (2);

[0094] Method (1):

[0095] The positions of non-cooperative space targets in Earth's inertial coordinate system are inferred based on their catalog information (the catalog information comes from ground-based sensing equipment and is obtained through long-term observation of non-cooperative space targets, including the six orbital parameters of the non-cooperative targets). Combined with the positions of on-orbit spacecraft in Earth's inertial coordinate system, the relative orbital information between the on-orbit spacecraft and each non-cooperative space target is calculated.

[0096] Based on the relative orbital information between the on-orbit spacecraft and various space non-cooperative targets, the space non-cooperative targets that have a collision risk with the on-orbit spacecraft are determined. The determined space non-cooperative targets are the space non-cooperative targets that require orbit transfer missions.

[0097] Method (2):

[0098] The relative position information between each non-cooperative space target and the on-orbit spacecraft is measured using the sensing mechanism installed on the on-orbit spacecraft;

[0099] According to the relative orbital information between the on-orbit spacecraft and various space non-cooperative targets, the space non-cooperative targets that have a collision risk with the on-orbit spacecraft are determined. The determined space non-cooperative targets are the space non-cooperative targets that need to perform orbit transfer missions.

[0100] Other steps and parameters are the same as those in the seventh embodiment.

[0101] Based on the relative orbital information between the on-orbit spacecraft and the space non-cooperative targets, it is possible to predict which space non-cooperative targets pose a threat to the on-orbit spacecraft. The space non-cooperative targets that pose a threat are the space non-cooperative targets that need to perform orbit transfer missions.

[0102] Specific embodiment 9: This embodiment differs from specific embodiments 7 or 8 in that, in step S2, the tether release module on the on-orbit spacecraft is used to control the tension on the tether to adjust the release speed of the tethered orbit transfer device, and the vector control device is used to provide the system with torque in and out of the orbital plane to control the tethered orbit transfer device to operate according to the calculated orbital maneuver trajectory. The specific process is as follows:

[0103] Step S21: The model of the space tether system consisting of the on-orbit spacecraft, tether, and tether orbit transfer device is constructed as follows:

[0104]

[0105] Where q = (θ, θ, φ) T , λ is the normalized parameter of the system rope length, is the first derivative of q, is the second derivative of q, θ is the orbital plane angle of the system, φ is the orbital plane angle of the system, u=(u l ,u θ ,u φ ) T ,u l is the normalized control variable of the tether tension, u θ is the normalized control quantity of the torque in the orbital plane of the system, u φ is the normalized control quantity of the system's out-of-plane torque, is the first-order coefficient matrix of the system, G(q) is the zero-order coefficient matrix of the system, and H(q) is the controller input coefficient matrix;

[0106] Both G(q) and H(q) change at any time with the state of the system; G(q) and H(q) are:

[0107]

[0108] Where e is the eccentricity of the Keplerian orbit of the system, v is the true approximate angle of the system, which represents the angle the system has rotated relative to the vernal equinox, k is the orbital coefficient, whose value is k = 1 + ecosv, and ρ is the bias coefficient;

[0109] Step S22: Design a fixed-time sliding surface structure based on the spatial tether system model:

[0110]

[0111] Among them, s i (q i ) is the sliding surface, q1=λ-1, q2=θ, q3=φ, It is q i The first derivative of |q i | isq i The absolute value of sgn(·) is the sign function, k i1 ,k i2 ,α i1 ,γ i1 is the sliding surface coefficient, and the sliding surface coefficient satisfies k i1 >0,k i2 >0,0<γ i1 <1<α i1 ;

[0112] Then the fixed-time controller output is:

[0113]

[0114] in, is a matrix The i-th row element, G i (q) is the i-th row element of G(q), H(q) ii -1 refers to H(q) -1 The element in row i and column i in , H(q) -1 is the inverse matrix of H(q), k i3 ,k i4 ,α i2 is the controller coefficient, and satisfies k i3 >0,k i4 >0,α i2 >1;

[0115] Step S23 , the tethered track transfer device release module and the vector control device control the tethered track transfer device to reach the location of the non-cooperative target along the track maneuvering trajectory according to the output of the fixed time controller.

[0116] Other steps and parameters are the same as those in the seventh or eighth embodiment.

[0117] Sensing systems installed on in-orbit spacecraft include, but are not limited to, optical imaging equipment, infrared, and radar. Based on the relative positional information between an in-orbit spacecraft and non-cooperative space targets, it is possible to predict which non-cooperative space targets pose a threat to the in-orbit spacecraft. These threatening non-cooperative space targets are those requiring orbital diversion missions.

[0118] This embodiment controls the release of the tethered orbital transfer device based on the initial release speed, initial angle, and orbital maneuvering trajectory after release obtained by motion planning. First, the release angle of the orbital transfer device is adjusted by using the steering adjustment unit installed on the on-orbit spacecraft. Then, the release speed is adjusted by obtaining a mechanism (which can be momentum transfer through collision, or the energy stored in the energy storage device can be released to the orbital transfer device by limiting the fuse) through the release speed adjustment module installed on the on-orbit spacecraft. The orbital transfer device will move away from the on-orbit spacecraft under the action of the initial speed. In the process of the orbital transfer device moving away from the on-orbit spacecraft, the tether release module on the on-orbit spacecraft can adjust the release speed of the orbital transfer device by controlling the tension on the tether. At the same time, the vector control device on the orbital transfer device can provide the system composed of the on-orbit spacecraft-tether-tethered orbital transfer device with torque in the orbital plane and out of the orbital plane, thereby adjusting the in-plane and out-of-plane angles of the system, and then controlling the orbital transfer device to reach the location of the non-cooperative target according to the preset orbit.

[0119] The fixed-time controller designed in this invention is applied to the release process of the space tether system, which can make the system reach the specified target point within a fixed time, and the stabilization time satisfies:

[0120]

[0121] Specific embodiment 10: This embodiment differs from any one of specific embodiments 7 to 9 in that the specific process of step S4 is as follows:

[0122] Step S41: using the electromagnetic thruster on the tethered track transfer device to eject high-speed charged ions to apply positive thrust to the tethered track transfer device;

[0123] Step S42: When the tethered track transfer device pulled by the conductive tether 8 is in the recovery process, the conductive tether 8 will cut the Earth's magnetic field, that is, electrons will accumulate on the conductive tether 8. The accumulated electrons are then emitted by the Spindt cathode emitter installed on the tethered track transfer device to form an electric current;

[0124] The current in step S43 and step S42 moves in the Earth's magnetic field to generate a Lorentz force, and the tethered track transfer device is braked using the Lorentz force and the forward thrust in step S41;

[0125] Step S44: Utilize the tether release module on the on-orbit spacecraft to recover the conductive tether 8, and pull the tether orbit transfer device to recover it to the on-orbit spacecraft.

[0126] The other steps and parameters are the same as those in any one of the seventh to ninth embodiments.

[0127] The above examples are merely illustrative of the calculation model and process of the present invention and are not intended to limit the embodiments of the present invention. Persons skilled in the art will readily appreciate that other variations or modifications based on the above description are possible. This list of embodiments is not exhaustive; however, any obvious variations or modifications derived from the technical solution of the present invention remain within the scope of protection of the present invention.

Claims

1. A space non-cooperative target orbit transfer system based on tether momentum isolation, characterized in that: The space non-cooperative target orbit transfer system comprises an on-orbit spacecraft and a tethered orbit transfer device, wherein the on-orbit spacecraft and the tethered orbit transfer device are connected via a conductive tether (8); The tethered track transfer device comprises a strike plate (1), an energy storage structure, a photoelectric sensor (4), a tethered track transfer device body (6), a Spindt cathode emitter (7), and a vector thruster (10); After the conductive tether (8) is inserted into the interior of the tethered track transfer device body (6), it is connected to the Spindt cathode emitter (7) on the surface of the tethered track transfer device body (6); The rope-tethered track transfer device body (6) is provided with an impact plate (1), a photoelectric sensor (4) and a vector thruster (10), and an energy storage structure is provided at a symmetrical position on the opposite surface of the impact plate (1) and the rope-tethered track transfer device body (6); The energy storage structure comprises a sleeve (2), a spring (3) and an electromagnet; The sleeve (2) is symmetrically arranged on the opposite surfaces of the rope track transfer device body (6) and the impact plate (1), and a hollow electromagnet is arranged at the non-fixed end of the sleeve (2); After the spring (3) passes through the symmetrically arranged sleeve (2) and the electromagnet, one end of the spring (3) is fixedly connected to the rope track transfer device body (6), and the other end is fixedly connected to the impact plate (1).

2. The space non-cooperative target orbit transfer system based on tether momentum isolation according to claim 1 is characterized in that: Four energy storage mechanisms are provided between the rope-tethered track transfer device body (6) and the impact plate (1).

3. The space non-cooperative target orbit transfer system based on tether momentum isolation according to claim 1 is characterized in that: The impact plate (1) is a circular plate with a partially hollowed-out area, and a photoelectric sensor (4) is provided on the upper surface of the impact plate (1); The photoelectric sensors (4) on the rope-tethered track transfer device body (6) and the impact plate (1) are used to observe and sense the relative position of the non-cooperative target and the rope-tethered track transfer device.

4. The space non-cooperative target orbit transfer system based on tether momentum isolation according to claim 1 is characterized in that: The rope-tethered track transfer device further comprises a first foldable solar wing (5) and a second foldable solar wing (9), and the first foldable solar wing (5) and the second foldable solar wing (9) are hinged on the rope-tethered track transfer device body (6).

5. The space non-cooperative target orbit transfer system based on tether momentum isolation according to claim 1 is characterized in that: The on-orbit spacecraft is provided with a tether release module, a release speed adjustment module and a steering adjustment unit; The conductive tether (8) is connected to a tether release module, the tether release module is used to release the tether track transfer device, the release speed adjustment module is used to adjust the speed of the track transfer device release, and the steering adjustment unit is used to adjust the angle of the track transfer device release.

6. The space non-cooperative target orbit transfer system based on tether momentum isolation according to claim 1, characterized in that: The rope-tethered track transfer device is provided with a vector control device, and the vector control device is used to provide the rope-tethered track transfer device with torque within the track plane and outside the track plane.

7. A space non-cooperative target orbit transfer method based on a space non-cooperative target orbit transfer system based on tether momentum isolation according to claim 1, characterized in that: The method specifically comprises the following steps: Step S1: Select a non-cooperative space target that needs to perform an orbit transfer mission based on the relative orbit information between the on-orbit spacecraft and the non-cooperative space target; then plan and calculate the initial velocity and initial angle of the tethered orbit transfer device after release, and the orbital maneuver trajectory of the tethered orbit transfer device after release; Step S2: Using the steering adjustment unit and release speed adjustment module on the on-orbit spacecraft to adjust the release of the tethered orbit transfer device according to the initial angle and initial speed. When the orbit transfer device is moving away from the on-orbit spacecraft, the tether release module on the on-orbit spacecraft controls the tension on the tether to adjust the release speed of the orbit transfer device. Using the vector control device, the system is provided with torque in and out of the orbital plane, thereby adjusting the system's in-plane and out-of-plane angles and controlling the orbit transfer device to operate according to the calculated orbital maneuver trajectory. Step S3: Using the photoelectric sensor on the tethered track transfer device to detect the position of the non-cooperative target, and then using the vector thruster (10) on the track transfer device to adjust the track transfer device posture so that the impact plate (1) faces the non-cooperative target, and then using the electromagnetic catapult collision device to send a release signal. After the spring is released, the impact plate (1) is used to impact the non-cooperative target; Step S4: Recover the conductive tether (8), and pull the tether track transfer device to recover it to the on-orbit spacecraft.

8. The space non-cooperative target orbit transfer method based on tether momentum isolation of a space non-cooperative target orbit transfer system according to claim 7, characterized in that: The method comprises selecting a non-cooperative space target that needs to perform an orbit transfer mission based on the relative orbit information between the on-orbit spacecraft and the non-cooperative space target; the specific method adopts the following method (1) or method (2); Method (1): The positions of non-cooperative space targets in the Earth's inertial coordinate system are inferred based on their catalog information. Combined with the positions of on-orbit spacecraft in the Earth's inertial coordinate system, the relative orbital information between the on-orbit spacecraft and each non-cooperative space target is calculated. Based on the relative orbital information between the on-orbit spacecraft and various space non-cooperative targets, the space non-cooperative targets that have a collision risk with the on-orbit spacecraft are determined. The determined space non-cooperative targets are the space non-cooperative targets that require orbit transfer missions. Method (2): The relative position information between each non-cooperative space target and the on-orbit spacecraft is measured using the sensing mechanism installed on the on-orbit spacecraft; According to the relative orbital information between the on-orbit spacecraft and various space non-cooperative targets, the space non-cooperative targets that have a collision risk with the on-orbit spacecraft are determined. The determined space non-cooperative targets are the space non-cooperative targets that need to perform orbit transfer missions.

9. The space non-cooperative target orbit transfer method based on tether momentum isolation of the space non-cooperative target orbit transfer system according to claim 8, characterized in that: In step S2, the tether release module on the on-orbit spacecraft is used to control the tension on the tether to adjust the release speed of the tethered orbit transfer device. The vector control device is used to provide the system with torque in and out of the orbital plane to control the tethered orbit transfer device to operate according to the calculated orbital maneuver trajectory. The specific process is as follows: Step S21: The model of the space tether system consisting of the on-orbit spacecraft, tether, and tether orbit transfer device is constructed as follows: Where q = (λ,θ,φ) T , λ is the normalized parameter of the system rope length, is the first derivative of q, is the second derivative of q, θ is the orbital plane angle of the system, φ is the orbital plane angle of the system, u=(u l ,u θ ,u φ ) T ,u l is the normalized control variable of the tether tension, u θ is the normalized control quantity of the torque in the orbital plane of the system, u φ is the normalized control quantity of the system's out-of-plane torque, is the first-order coefficient matrix of the system, G(q) is the zero-order coefficient matrix of the system, and H(q) is the controller input coefficient matrix; G(q) and H(q) are: Where e is the eccentricity of the Keplerian orbit of the system, v is the true near angle of the system, k is the orbit coefficient, and ρ is the bias coefficient; Step S22: Design the sliding surface structure according to the spatial tether system model: Among them, s i (q i ) is the sliding surface, q1=λ-1, q2=θ, q3=φ, It is q i The first derivative of |q i | isq i The absolute value of sgn(·) is the sign function, k i1 ,k i2 ,α i1 ,γ i1 is the sliding surface coefficient, and the sliding surface coefficient satisfies k i1 >0,k i2 >0,0<γ i1 <1<α i1 ; The controller output is: in, is a matrix The i-th row element, G i (q) is the i-th row element of G(q), H(q) ii -1 refers to H(q) -1 The element in row i and column i in , H(q) -1 is the inverse matrix of H(q), k i3 ,k i4 ,α i2 is the controller coefficient, and satisfies k i3 >0,k i4 >0,α i2 >1; Step S23 : The tethered release module and the vector control device control the tethered track transfer device to reach the location of the non-cooperative target along the track maneuvering trajectory according to the output of the controller.

10. The space non-cooperative target orbit transfer method based on tether momentum isolation of a space non-cooperative target orbit transfer system according to claim 9, characterized in that: The specific process of step S4 is as follows: Step S41: using the electromagnetic thruster on the tethered track transfer device to eject high-speed charged ions to apply positive thrust to the tethered track transfer device; Step S42: When the tethered track transfer device pulled by the conductive tether (8) is in the recovery process, the conductive tether (8) will cut the earth's magnetic field, that is, electrons will be gathered on the conductive tether (8), and then the Spindt cathode emitter installed on the tethered track transfer device will emit the gathered electrons to form an electric current; The current in step S43 and step S42 moves in the Earth's magnetic field to generate a Lorentz force, and the tethered track transfer device is braked using the Lorentz force and the forward thrust in step S41; Step S44: Utilize the tether release module on the on-orbit spacecraft to recover the conductive tether (8), and pull the tether track transfer device to recover it to the on-orbit spacecraft.

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