Space-based interception orbital transfer strategy solving method applied to different-plane non-circular orbits

Abstract

The invention discloses a space-based interception orbital transfer strategy solving method applied to a different-plane non-circular orbit, and is applied to the field of space attack and defense. The method comprises the following specific steps: determining initial orbit parameters of a task spacecraft and a target spacecraft; determining orbits of the task spacecraft and the target spacecraft based on the initial orbit parameters of the task spacecraft and the target spacecraft; solving an interception moment and an interception position coordinate of the task spacecraft; according to the technical scheme, for a non-planar non-circular orbit spacecraft interception scene, the problem that a task spacecraft intercepts a target spacecraft within specified constraint time is solved, and the mode of staged decomposition of an interception task, staged solving and multi-stage constraint condition fusion is adopted; the method at least has the advantages that complex tasks can be simply disassembled in a staged mode, and space task planning requirements can be met.

Description

Technical Field

[0001] The present application relates to the field of space attack and defense, and specifically to a method for solving a space-based interception and trajectory change strategy applied to unequal-planar non-circular orbits. Background Art

[0002] With the advancement and development of aerospace technology, the pattern of future wars will also extend from land, sea and air operations to space operations. The development of space attack and defense systems, such as intercepting or directly destroying enemy targets, prevents the enemy from formulating corresponding war strategies based on effective real-time information, thereby greatly weakening the enemy's combat effectiveness. In order to better grasp the initiative of future wars, countries have begun to formulate and adopt a series of development strategies.

[0003] Spacecraft interception can be carried out by space-based and ground-based launches, among which space-based is divided into satellites or other spacecraft, and ground-based is divided into surface / underwater, land, and near-ground air. Among them, space-based interceptors are the key research objects and directed energy spacecraft of this invention. Kinetic weapons use kinetic spacecraft to directly collide to achieve target interception, and can be deployed on ships, ground, space shuttles, or spacecraft. Directed energy anti-satellite weapons mainly destroy sensitive components or satellite structures of target spacecraft, which can be achieved by launching high-power microwave beams, particle beams or high-energy lasers from space platforms, air or ground.

[0004] Research on orbit change strategies is mostly conducted on space targets in the same orbital plane, while the Lambert orbit change method is mostly used for space targets in different orbital planes. Since tracking the target requires time and distance, it usually requires the service vehicle to carry a large amount of orbit change fuel. The tracking time, speed and encounter position of the target spacecraft in the orbit change strategy have a huge impact on the fuel carried. In practice, research institutions and others have done very little research on spacecraft orbit change strategies for special scenarios such as different orbital planes.

[0005] Therefore, a new solution is needed to solve the problem of determining the trajectory change interception strategy in the scenario where the interceptor and the intercept target are in non-circular orbits, and to save fuel by quickly and accurately intercepting and maneuvering to correct the trajectory change speed. Summary of the invention

[0006] In view of this, the technical solution of the present invention provides a method for solving a space-based interception and trajectory change strategy applied to a non-coplanar non-circular orbit, comprising:

[0007] Determine the initial orbital parameters of the mission spacecraft and the target spacecraft;

[0008] Determining the orbits of the mission spacecraft and the target spacecraft based on the initial orbital parameters of the mission spacecraft and the target spacecraft;

[0009] Solve the interception time and interception position coordinates of the mission spacecraft;

[0010] Based on the orbit of the mission spacecraft, solving the orbital plane normal vector;

[0011] Solving the intersection coordinates of the orbital planes of the target spacecraft and the mission spacecraft based on the golden section algorithm;

[0012] Solving the time and position coordinates corresponding to the arrival of the target spacecraft at the intersection coordinates;

[0013] Determine the orbital pulse strategy for the mission spacecraft to reach the interception position at the interception time;

[0014] Based on the interception position coordinates, a Hohmann orbit change strategy is adopted; the orbit is changed at the perigee; based on the Kepler orbit algorithm, the coordinate position of the mission spacecraft at the perigee is solved;

[0015] Determine that the orbit of the mission spacecraft after the Hohmann orbit change at perigee is the transfer orbit, and the mission spacecraft will pass the interception point in the transfer orbit;

[0016] According to algorithm A, solving the transfer trajectory;

[0017] Based on the transfer orbit, solve the time difference from the mission spacecraft to the perigee and the interception point;

[0018] Solve the orbit pulse time, pulse position coordinates and pulse velocity increment;

[0019] Determine the optimal interception strategy.

[0020] Furthermore, the initial orbital parameters of the mission spacecraft and the target spacecraft are: orbital semi-major axis, orbital eccentricity, orbital inclination, right ascension of ascending node, argument of perigee, and true anomaly.

[0021] Furthermore, in Kepler's algorithm, R = [a(1-e ^ 2)] / (1+e cos(V))

[0022] Where R is the distance from the point to the center of the sphere; a is the semi-major axis; e is the eccentricity; V is the true anomaly angle. Assume a*(1-e ^ 2) is a constant, and there are points a and b in a certain orbit that satisfy the following algorithm A:

[0023]

[0024]

[0025] Furthermore, the algorithm A can quickly and accurately generate a transfer orbit passing through two designated points based on the orbital parameters of the mission spacecraft.

[0026] Furthermore, the solution method of the interception and trajectory change strategy is applied in the following scenario: the mission spacecraft and the target spacecraft are in unconstrained, skewed, non-circular orbits.

[0027] Furthermore, the out-of-plane interception mission of the mission spacecraft on the target spacecraft is subject to time constraints.

[0028] Further, solving the track surface normal vector based on the track parameters;

[0029] Furthermore, a mission intercept point solution algorithm is constructed based on the orbital plane normal vector and the golden section method, and the true anomaly angle of the target orbit intercept point that meets the mission requirements is quickly and iteratively solved.

[0030] Furthermore, by solving the true anomaly angle corresponding to the optimal solution of the golden section method, the Kepler algorithm is used to solve the position and time series of the interception point.

[0031] Furthermore, the interception problem is decomposed into solving two transfer orbits: transfer orbit δ and transfer orbit O.

[0032] Furthermore, the problem of solving the transfer orbit δ is converted into: solving the velocity vectors of multiple pulse excitations in the scenario of an interception mission of a spacecraft in a non-circular orbit; and correcting the interception and orbit transfer speed of the mission spacecraft by solving the velocity vectors of the multiple pulse maneuvers.

[0033] Compared with the prior art, the technical solution of the present invention aims at the interception scenario of spacecraft in unequal non-circular orbits, solves the problem of the mission spacecraft intercepting the target spacecraft within the specified constraint time, and adopts the method of decomposing the interception task into stages, solving the problem stage by stage, and fusing the multi-stage constraint conditions; the beneficial effects that can be achieved include at least: it can simply disassemble the complex task into stages, provide a solution method and layout strategy for the spacecraft interception problem in unequal non-circular orbits, and support the needs of space mission planning; it can quickly and accurately generate a transfer orbit passing through two specified points, and solve the interception problem of space missions in unequal non-circular orbits; it can save fuel, and all the transfer modes are tangential maneuvers, ensuring that the mission spacecraft completes the transfer and change of orbit at the lowest speed and with the least fuel. BRIEF DESCRIPTION OF THE DRAWINGS

[0034] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0035] Figure 1 It is a schematic diagram of the interception scenario between the mission spacecraft and the target spacecraft. DETAILED DESCRIPTION

[0036] The embodiments of the present application are described in detail below with reference to the accompanying drawings.

[0037] The following describes the implementation methods of the present application through specific examples, and those skilled in the art can easily understand other advantages and effects of the present application from the contents disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. The present application can also be implemented or applied through other different specific implementation methods, and the details in this specification can also be modified or changed in various ways based on different viewpoints and applications without departing from the spirit of the present application. It should be noted that, in the absence of conflict, the following embodiments and the features in the embodiments can be combined with each other. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in the field without making creative work belong to the scope of protection of the present application.

[0038] It should be noted that various aspects of the embodiments within the scope of the appended claims are described below. It should be apparent that the aspects described herein can be embodied in a wide variety of forms, and any specific structure and / or function described herein is merely illustrative. Based on the present application, it should be understood by those skilled in the art that an aspect described herein can be implemented independently of any other aspect, and two or more of these aspects can be combined in various ways. For example, any number and aspect described herein can be used to implement the device and / or practice the method. In addition, other structures and / or functionalities other than one or more of the aspects described herein can be used to implement this device and / or practice this method.

[0039] It should also be noted that the illustrations provided in the following embodiments are only schematic illustrations of the basic concept of the present application. The drawings only show components related to the present application rather than being drawn according to the number, shape and size of components in actual implementation. In actual implementation, the type, quantity and proportion of each component may be changed arbitrarily, and the component layout may also be more complicated.

[0040] Additionally, in the following description, specific details are provided to facilitate a thorough understanding of the examples. However, it will be understood by those skilled in the art that the examples can be practiced without these specific details.

[0041] The present invention discloses a method for determining a strategy for space-based interception of targets in skew non-circular orbits. For space attack and defense scenarios, consider that the mission spacecraft and the target spacecraft are in skew orbital planes; the mission spacecraft and the target spacecraft are in universal orbits, that is, the orbital eccentricity range of their respective orbits satisfies 0≤e<1; the mission of the mission spacecraft is to intercept the target spacecraft through the Hohmann orbit change strategy within a specified constraint time. The present invention discloses a method for solving a mission spacecraft long-range pulse interception orbit maneuvering strategy for the above-mentioned application scenario requirements, so that the mission spacecraft can achieve point-to-point interception of the target spacecraft within a specified constraint time.

[0042] The technical solution of the present invention is explained below with an example.

[0043] like Figure 1 As shown, OrbitB represents the target spacecraft; OrbitA1 represents the initial orbit of the mission spacecraft; OrbitA2 represents the transfer orbit δ as described in step 11 of the invention content, and OrbitA3 represents the transfer orbit as described in step 7 of the invention content. The example scenario requires the mission spacecraft to transfer orbits from OrbitA1 to OrbitA2 and OrbitA3, and intercept the target spacecraft at the Pimpact point at the corresponding time. The method and steps for solving the maneuvering strategy to achieve the mission are as follows:

[0044] (1) Phase 1: Determine the initial orbital parameters of the mission spacecraft and the target spacecraft, including:

[0045] Step 1: Determine the six orbital elements of OrbitA1;

[0046] Step 2: Determine the six orbital elements of OrbitB;

[0047] (2) Phase II: Determine the interception time and position of the mission spacecraft, including:

[0048] Step 3: According to the orbits of OrbitA1 and OrbitB determined in step 1, solve the normal vector of the orbital plane;

[0049] Step 3.1: Calculate the space coordinate Pa of the mission spacecraft when the true anomaly v = 0;

[0050] Step 3.2: Calculate the space coordinate Pb of the mission spacecraft when the true anomaly V = x, x ≠ 0;

[0051] Step 3.3: Based on the principle of vector cross multiplication, solve the cross product of the above two coordinates This result is the normal vector of the orbital plane;

[0052] Step 4: Calculate the intersection of the orbital planes of the target spacecraft and the mission spacecraft based on the golden section algorithm;

[0053] Step 4.1: Take the true anomaly angle v in the target spacecraft orbital parameters as the optimized variable;

[0054] Step 4.2: Take the distance between the orbital position of the target spacecraft and the orbital plane of the mission spacecraft as the objective function;

[0055] Solution 4.3: Using the above steps 4.1 and 4.2 as algorithm conditions, find the optimal solution of the golden section algorithm;

[0056] Solution 4.4: If the tolerance value of the golden section method is set small enough, the result of the golden section method can be approximately considered to be the global optimal solution, and the error generated can be ignored;

[0057] Step 5: Calculate the time point and coordinates corresponding to the target spacecraft's arrival at the intersection coordinates;

[0058] Step 5.1: Based on the optimization result output by the golden section method solved in step 4, solve the spatial coordinate Pimpact of the result as the first interception point;

[0059] Step 5.2: Based on the optimization result output by the golden section method solved in step 4, increase its phase by 180 degrees, and solve the spatial coordinate Pimpact2 of the result as the second interception point;

[0060] Step 5.3: Based on the optimization result output by the golden section method solved in step 5.1, solve the time Timpact corresponding to the result. The solved time result is not unique;

[0061] Step 5.4: Based on the optimization result output by the golden section method solved in step 5.2, solve the time Timpact2 corresponding to the result. The solved time result is not unique;

[0062] (3) Phase 3: Solve the orbital pulse strategy for the mission spacecraft to reach the specified interception position solved in Phase 2 at the specified time solved in Phase 2, including:

[0063] Step 6: Take an intercept point solved in stage 2 and temporarily count it as Pimpact;

[0064] Step 6: Considering that the mission spacecraft adopts the Hohmann orbit change strategy and changes its orbit at the perigee Pa, the coordinate position Pa of the mission spacecraft at the perigee is solved based on the Kepler orbit algorithm;

[0065] Step 7: Consider that the orbit of the mission spacecraft after the Hohmann orbit change at the perigee Pa point is the transfer orbit (OrbitA3), and the mission spacecraft will pass through the Pimpact point on the transfer orbit;

[0066] Step 8: Solve the transfer orbit described in Step 7 by:

[0067] Step 8.1: To save fuel, consider that only tangential maneuvers are performed in the transfer orbit, and the orbital inclination, perigee argument, and ascending node right ascension in the orbital parameters remain unchanged;

[0068] Step 8.2: Solve for Vdeg = ∠P a OP impact , that is, when the mission spacecraft is at the Pimpact point on the transfer orbit, the corresponding true anomaly angle is Vdeg;

[0069] Step 8.3: Based on the above parameters, and Figure 1 Algorithm A shown here solves the semi-major axis and eccentricity parameters of the transfer orbit (OrbitA3);

[0070] Step 8.4: Based on the processes in step 8 above, the orbit semi-major axis, eccentricity, orbit inclination, argument of perigee, right ascension of ascending node, and true anomaly angle can be constructed to construct the transfer orbit O: OrbitA3;

[0071] Step 9: Based on the transfer orbit parameters solved in step 8, calculate the time consumption Ttransfer of the mission spacecraft from point Pa to Pimpact;

[0072] Step 9.1: The mean anomaly angle corresponding to point Pa is 0;

[0073] Step 9.2: The true anomaly angle corresponding to the Pimpact point is Vdeg, based on

[0074] M=Ee·sin(E)

[0075]

[0076] The mean anomaly angle M of Pimpact can be solved;

[0077] Step 9.3: Based on

[0078]

[0079] n·ΔT=ΔM

[0080] The shortest transfer time Ttransfer of the mission spacecraft from point Pa to point Pimpact on the transfer orbit can be solved;

[0081] Step 9.4: Based on

[0082] n·T=2π

[0083] The orbital period T of the transfer orbit can be solved.

[0084] Step 9.5: Transfer time Ttransfer = Ttransfer + k·T, where k is a non-negative integer;

[0085] Step 10: Based on the above problem description and parameter solution, the final interception problem is converted to: solve the mission spacecraft through the Hohmann orbit change to ensure that it arrives at the perigee Pa on time from the current time (deltaT = Timpact-Now-Ttransfer), and deltaT meets the time constraint. Now in the formula represents the time value corresponding to the current moment;

[0086] Step 11: Solve the Hohmann transfer orbit (OrbitA2);

[0087] Step 11.1: There are many ways to implement the Hohmann track change strategy, and the present invention only selects one of them as a solution;

[0088] Step 11.2: Consider that the Hohmann orbit change always occurs at point Pa and the orbit change is only once. The orbit period Period after the orbit change satisfies:

[0089] k*Period+Toffset=deltaT

[0090] Where k is a positive integer, Toffset is the time when the mission spacecraft reaches point Pa next time;

[0091] Step 11.3: Starting from k = 1, calculate the orbital period Period under various k values.

[0092]

[0093] Solve for the orbital semi-major axis a;

[0094] Step 11.4: Based on the steps in step 11 above, a series of result data sets can be obtained;

[0095] Step 11.5: From the result data set, take the one whose orbital semi-major axis is closest to the current orbital semi-major axis of the mission spacecraft, which is regarded as the final solution, and record the orbital semi-major axis a′ at this time;

[0096] Step 11.6: Based on the orbit semi-major axis a′ and the coordinates of point Pa obtained in step 11.5, calculate the orbital eccentricity e′ of the orbit;

[0097] Step 11.6: Construct the Hohmann transfer orbit δ, i.e., OrbitA2, using the semi-major axis a′, eccentricity e, inclination of the mission spacecraft’s current orbit, argument of perigee, and right ascension of the ascending node as orbital parameters;

[0098] Step 12: Solve the orbit pulse time, position, pulse velocity increment, etc., including:

[0099] Step 12.1: Starting from the current moment, based on

[0100]

[0101] Calculate the time Toffset from the next arrival of the mission spacecraft at point Pa;

[0102] Step 12.2: Based on the current orbital parameters of the mission spacecraft, calculate the speed at which the mission spacecraft will reach point Pa next time Parameters such as time;

[0103] Step 12.3: Based on the Hohmann transfer orbit δ, solve for the velocity of the mission spacecraft at point Pa

[0104] Parameters such as time;

[0105] Step 12.4: Based on the Hohmann transfer orbit O, solve for the velocity of the mission spacecraft at point Pa

[0106] Parameters such as time;

[0107] Step 12.5: Based on the parameters solved in step 12 above, solve the first pulse velocity increment as: The second pulse velocity increment is solved as:

[0108] Each embodiment in this specification is described in a progressive manner, and the same or similar parts between the embodiments can be referred to each other, and each embodiment focuses on the differences from other embodiments. In particular, for the product embodiment described later, since it corresponds to the method, the description is relatively simple, and the relevant parts can be referred to the partial description of the system embodiment.

[0109] The above is only a specific implementation of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed in the present application should be included in the protection scope of the present application. Therefore, the protection scope of the present application shall be based on the protection scope of the claims.

Claims

1. A method for solving a space-based interception and trajectory change strategy for non-coplanar non-circular orbits, characterized in that: include: Determine the initial orbital parameters of the mission spacecraft and the target spacecraft; Determining the orbits of the mission spacecraft and the target spacecraft based on the initial orbital parameters of the mission spacecraft and the target spacecraft; Solve the interception time and interception position coordinates of the mission spacecraft; Based on the orbit of the mission spacecraft, solving the orbital plane normal vector; Solving the intersection coordinates of the orbital planes of the target spacecraft and the mission spacecraft based on the golden section algorithm; Solving the time and position coordinates corresponding to the arrival of the target spacecraft at the intersection coordinates; Determine the orbital pulse strategy for the mission spacecraft to reach the interception position at the interception time; Based on the interception position coordinates, a Hohmann orbit change strategy is adopted; the orbit is changed at the perigee; based on the Kepler orbit algorithm, the coordinate position of the mission spacecraft at the perigee is solved; Determine that the orbit of the mission spacecraft after the Hohmann orbit change at perigee is the transfer orbit, and the mission spacecraft will pass the interception point in the transfer orbit; According to algorithm A, solving the transfer trajectory; Based on the transfer orbit, solve the time difference from the mission spacecraft to the perigee and the interception point; Solve the orbital pulse time, pulse position coordinates and pulse velocity increment.

2. The method for solving the space-based interception and orbit change strategy applied to non-coplanar non-circular orbits according to claim 1 is characterized in that: The initial orbital parameters of the mission spacecraft and the target spacecraft are: orbital semi-major axis, orbital eccentricity, orbital inclination, right ascension of ascending node, argument of perigee, and true anomaly.

3. The method for solving the space-based interception and orbit change strategy applied to non-coplanar non-circular orbits according to claim 1 is characterized in that: Kepler's algorithm exists: R = [a(1-e^2)] / (1+e cos(V)); Among them, R is the distance from the point to the center of the sphere; a is the semi-major axis; e is the eccentricity; V is the true anomaly angle R = [a*(1-e^2)] / (1+e*cos(V)) =>a*(1-e^2)=R*(1+e*cos(V)) Assume that there are points a and b that satisfy the following algorithm A: Ra*(1+e*cos(Va))=Rb*(1+e*cos(Vb)) =>e=(Ra-Rb) / (Rb*cos(Vb)-Ra*cos(Va)) =>a=Ra*(1+e*cos(Va)) / (1-e^2).

4. The method for solving the space-based interception and orbit change strategy applied to non-coplanar non-circular orbits according to claims 1 to 3 is characterized in that: The algorithm A can quickly and accurately generate a transfer orbit passing through two designated points based on the orbital parameters of the mission spacecraft.

5. The method for solving the space-based interception and orbit change strategy applied to non-coplanar non-circular orbits according to claims 1 to 3 is characterized in that: The solution method of the interception and trajectory change strategy is applied in the scenario where the mission spacecraft and the target spacecraft are in unconstrained skewed non-circular orbits.

6. The method for solving the space-based interception and orbit change strategy applied to non-coplanar non-circular orbits according to claims 1 to 3 is characterized in that: The out-of-plane interception mission of the mission spacecraft on the target spacecraft is subject to time constraints.

7. The method for solving the space-based interception and orbit change strategy applied to non-coplanar non-circular orbits according to claims 1 to 3 is characterized in that: The orbital plane normal vector is solved based on the orbital parameters.

8. The method for solving the space-based interception and orbit change strategy applied to non-coplanar non-circular orbits according to claims 1 to 3 is characterized in that: The mission interception point solution algorithm is constructed based on the orbital plane normal vector and the golden section method, and the true anomaly angle of the target orbit interception point that meets the mission requirements is solved by rapid iteration.

9. The method for solving the space-based interception and orbit change strategy applied to non-coplanar non-circular orbits according to claims 1 to 3, characterized in that: By solving the true anomaly angle corresponding to the optimal solution of the golden section method, the Kepler algorithm is used to solve the position and time series of the interception point.

10. According to the method for solving the space-based interception and orbit change strategy applied to eccentric non-circular orbits as described in claims 1 to 3, the interception problem is decomposed into solving two transfer orbits: transfer orbit δ and transfer orbit θ.

11. The method for solving the space-based interception and orbit change strategy applied to non-coplanar non-circular orbits according to claims 1 to 10, characterized in that: The problem of solving the transfer orbit δ is converted into: solving the velocity vector of multiple pulse maneuvers in the scenario of interception mission of spacecraft in non-circular orbit; The interception trajectory change speed of the mission spacecraft is corrected by solving the velocity vectors of the multiple pulse maneuvers.

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