An improved closed loop guidance method for reusable rockets

Abstract

The application provides an improved closed-loop guidance method suitable for a reusable rocket, and relates to the technical field of rocket recovery guidance. The method comprises the following steps: determining the velocity and position in the launch inertial system at the engine shutdown time of the power deceleration stage according to the initial value of the remaining flight time of the active stage, the real-time gravity acceleration and the initial value of the program angle of the power deceleration stage; determining the absolute longitude difference between the position in the launch inertial system at the engine shutdown time of the power deceleration stage and the target point; determining the remaining flight time of the passive stage; and determining the required velocity and program angle calculation value of the closed-loop guidance. According to the application, the calculation method of predicting the state of the shutdown point can be used to calculate the required velocity, the program angle change rate of the power deceleration stage is reduced, the control effect is optimized, and the stability is improved. Moreover, the calculated required velocity can reach the position constraint and velocity direction constraint of the passive stage end point after the rocket is shut down, the structure is simple, the operation efficiency is high, and the method is convenient to run on different hardware.

Description

Technical Field

[0001] This invention relates to the field of rocket recovery guidance technology, and in particular to an improved closed-circuit guidance method suitable for reusable rockets. Background Technology

[0002] Reusable launch vehicles, by recovering and reusing the first stage, reduce the cost per launch and lessen the constraints of production on launch frequency, making them a crucial link in unleashing the potential of the space industry and supporting large-scale constellation deployment. The powered deceleration phase is the first flight phase after the first and second stages separate, where the flight trajectory is adjusted, and it is critical for range correction and precise control of the landing ignition timing. Among related technologies, closed-circuit guidance algorithms are a mature algorithm in the guidance field, with advantages such as simple algorithm structure and high computational efficiency, and are often used to control the terminal position of the passive phase after shutdown. However, this method still has significant limitations in solving the guidance problem of the powered deceleration phase. For example, the powered deceleration trajectory is not a minimum energy trajectory, which is inconsistent with the assumptions of traditional closed-circuit guidance, and some algorithm process parameters cannot be directly calculated analytically; the velocity in the powered deceleration phase is relatively small, resulting in rapid changes in the calculated gain velocity and a large rate of change in the output program angle, which is not conducive to stable control; it is difficult to constrain the velocity in the passive phase, and it is also difficult to control the dynamic pressure and trajectory inclination of the subsequent passive phase.

[0003] The information disclosed in the background section of this application is intended only to enhance the understanding of the general background of this application and should not be construed as an admission or in any way implying that the information constitutes prior art known to those skilled in the art. Summary of the Invention

[0004] This invention provides an improved closed-circuit guidance method suitable for reusable rockets, which can solve the technical problems of difficulty in analyzing and calculating process parameters, difficulty in achieving stable control, and difficulty in constraining the passive phase velocity in related technologies.

[0005] According to a first aspect of the present invention, an improved closed-circuit guidance method suitable for reusable rockets is provided, comprising: Based on the initial value of the remaining flight time in the active phase, the real-time gravitational acceleration, and the initial value of the programmed angle in the powered deceleration phase, determine the velocity and position in the launch inertial frame at the moment of shutdown in the powered deceleration phase; Based on the position of the target point in the launch inertial frame at the moment of shutdown of the power deceleration phase, and the position of the target point in the launch frame, determine the absolute longitude difference between the position of the target point in the launch inertial frame at the moment of shutdown of the power deceleration phase. The remaining flight time of the passive segment is determined based on the absolute longitude difference and the position of the target point under the launch system. Based on the remaining flight time of the passive phase, the absolute longitude difference, and the velocity in the launch inertial frame at the moment of shutdown in the powered deceleration phase, the required velocity and program angle calculation values ​​for closed-circuit guidance are determined.

[0006] According to the present invention, determining the velocity and position in the launch inertial frame at the moment of shutdown during the power deceleration phase includes: The remaining flight time of the active phase is determined based on the initial value of the remaining flight time of the active phase and the number of calculation cycles. Based on the real-time gravitational acceleration and the initial value of the program angle during the dynamic deceleration phase, the differential equation to be solved is set. Based on the remaining flight time of the active phase and the differential equations to be solved, determine the velocity and position in the launch inertial frame at the moment of shutdown during the powered deceleration phase.

[0007] According to the present invention, determining the absolute longitude difference between the launch position in the inertial frame and the target point at the moment of shutdown during the power deceleration phase includes: Obtain the geocentric radius vector of the origin of the launch inertial frame and the transformation matrix from the launch inertial frame to the launch frame; Based on the position in the launch inertial frame at the moment of shutdown during the power deceleration phase, the geocentric radius of the origin of the launch inertial frame, and the transformation matrix from the launch inertial frame to the launch frame, the geocentric radius of the predicted shutdown point is determined. The geocentric radius of the target point is determined based on the geocentric radius of the origin of the launch inertial frame and the position of the target point in the launch frame. Based on the geocentric radius vector of the predicted shutdown point and the geocentric radius vector of the target point, determine the geocentric latitude of the predicted shutdown point and the geocentric latitude of the target point; Based on the geocentric latitude of the predicted shutdown point, the geocentric latitude of the target point, the geographical latitude of the launch point, and the geocentric radius vector of the predicted shutdown point, determine the absolute longitude difference between the launch inertial frame position and the target point at the moment of shutdown during the powered deceleration phase.

[0008] According to the present invention, determining the remaining flight time of the passive segment includes: Based on the absolute longitude difference and the position of the target point in the launching system, set the constructor; Transform the constructor into an iterative function of iterative form; The iterative function is solved iteratively, and the remaining flight time of the passive segment is obtained if the iterative function meets the preset convergence condition.

[0009] According to the present invention, setting a constructor includes: Determine the range angle based on the absolute longitude difference; The trajectory inclination angle is determined based on the position of the target point in the launch frame and the transformation matrix from the launch frame to the launch inertial frame at the end of the passive phase. Determine the passive section trajectory semi-bore based on the range angle and the trajectory inclination angle; Based on the passive section trajectory semi-circle, ballistic inclination angle, and range angle, determine the predicted shutdown point and the apogee angle of the target point; The track parameters are determined based on the apogee angle and the semi-circular diameter of the passive track segment; Based on the orbital parameters, set the constructor.

[0010] According to the present invention, determining the required velocity and program angle values ​​for closed-circuit guidance includes: The required speed for closed-loop guidance is determined based on the remaining flight time and absolute longitude difference of the passive segment. The program angle calculation value is determined based on the required speed for closed-circuit guidance and the speed in the launch inertial frame at the moment of shutdown during the power deceleration phase.

[0011] According to the present invention, determining the required velocity for closed-circuit guidance includes: Based on the passive section track semi-circle, track parameters, and ballistic inclination, determine the required velocity modulus and predict the ballistic inclination in the shutdown state; The angle between the orbital plane and the meridian plane is determined based on the absolute longitude difference, the remaining flight time of the passive segment, and the range angle. The required velocity for closed-circuit guidance is determined based on the angle between the orbital plane and the meridional plane, the required velocity modulus, and the predicted trajectory inclination angle in the shutdown state.

[0012] According to a second aspect of the invention, an improved closed-circuit guidance system suitable for reusable rockets is provided. The velocity and position module determines the velocity and position in the launch inertial frame at the moment of shutdown during the powered deceleration phase, based on the initial value of the remaining flight time during the active phase, the real-time gravitational acceleration, and the initial value of the programmed angle during the powered deceleration phase. The absolute longitude difference module determines the absolute longitude difference between the position in the launch inertial frame at the moment the power deceleration phase is shut down and the position of the target point in the launch frame. The remaining flight time module determines the remaining flight time of the passive segment based on the absolute longitude difference and the position of the target point under the launch system; The required speed and program angle module determines the required speed and program angle values ​​for closed-circuit guidance based on the remaining flight time of the passive phase, the absolute longitude difference, and the velocity in the launch inertial frame at the moment of shutdown in the powered deceleration phase.

[0013] According to a third aspect of the present invention, an improved closed-circuit guidance device suitable for reusable rockets is provided, comprising: a processor; a memory for storing processor-executable instructions; wherein the processor is configured to invoke the instructions stored in the memory to execute the improved closed-circuit guidance method suitable for reusable rockets.

[0014] According to a fourth aspect of the present invention, a computer-readable storage medium is provided having computer program instructions stored thereon, which, when executed by a processor, implement the improved closed-circuit guidance method applicable to reusable rockets.

[0015] By adopting the above technical solution, the present invention can achieve the following technical effects: The improved closed-loop guidance method for reusable rockets proposed in this application employs a calculation method based on predicting the shutdown point state. This method can calculate the velocity required for closed-loop guidance and constrain the position and velocity direction of the passive phase endpoint after rocket shutdown. It is applicable to the powered deceleration phase of reusable rockets. The proposed solution uses a method based on predicting the shutdown point state to calculate the required velocity, reducing the rate of change of the programmed angle during powered deceleration, optimizing the control effect, and improving stability. Furthermore, it can solve the closed-loop guidance algorithm parameters under non-minimum energy trajectory constraints, enabling the engineering application of the closed-loop guidance algorithm in the powered deceleration phase during the first-stage recovery of a rocket. Moreover, this solution is not limited to the reentry deceleration phase of reusable rockets; it is applicable in scenarios requiring control of the passive phase endpoint after shutdown, such as high-altitude recovery phases or the rocket ascent phase. It achieves high guidance accuracy, has a simple algorithm structure, high computational efficiency, and is easy to run on different hardware.

[0016] The improved closed-circuit guidance method for reusable rockets proposed in this application can employ a calculation method to predict the shutdown point state. When determining the velocity and position in the launch inertial frame at the shutdown moment of the powered deceleration phase, the remaining flight time of each active phase can be calculated by using a pre-loaded value decreasing calculation cycle method. Furthermore, a neural network can be fitted based on the offline simulation results of a large sample size, and the remaining flight time can be calculated through the input flight state, thereby improving the accuracy of the prediction.

[0017] This application presents an improved closed-loop guidance method for reusable rockets, employing a calculation approach that predicts the shutdown point state. When setting the constructor function, Newton's iteration method is used to calculate nonlinear equations, enhancing the robustness and speed of the solution process. It can also combine the two-body problem to solve some of the equations analytically. This method expands the application scenarios of closed-loop guidance algorithms under non-minimum energy orbital conditions, enabling closed-loop guidance algorithms to meet the guidance requirements of the powered deceleration phase of reusable rockets.

[0018] Furthermore, the improved closed-circuit guidance method for reusable rockets in this application adopts a calculation method for predicting the shutdown point state. It can use a constant program angle assumption to predict the shutdown state during the powered deceleration phase, taking into account different application scenarios or flight trajectories. This improves the adaptability of the closed-circuit guidance algorithm to the powered deceleration phase. At the same time, it can be further improved based on different program angle assumption forms, thereby extending its application to other stages of rocket flight.

[0019] It should be understood that the foregoing general description and the following detailed description are exemplary and explanatory only, and are not intended to limit the invention. Other features and aspects of the invention will become clearer from the following detailed description of exemplary embodiments with reference to the accompanying drawings. Attached Figure Description

[0020] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other embodiments can be obtained based on these drawings without creative effort. Figure 1 An exemplary flowchart of an improved closed-circuit guidance method for reusable rockets according to an embodiment of the present invention is shown. Figure 2 A flowchart of an improved closed-circuit guidance method for reusable rockets according to an embodiment of the present invention is shown as an example. Figure 3 A block diagram of an improved closed-circuit guidance system for reusable rockets according to an embodiment of the present invention is shown as an example. Detailed Implementation

[0021] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0022] The technical solution of the present invention will be described in detail below with reference to specific embodiments. These specific embodiments can be combined with each other, and the same or similar concepts or processes may not be described again in some embodiments.

[0023] Figure 1 An exemplary flowchart of an improved closed-loop guidance method for reusable rockets according to an embodiment of the present invention is shown, the method comprising: Step S1: Based on the initial value of the remaining flight time in the active phase, the real-time gravitational acceleration, and the initial value of the program angle in the powered deceleration phase, determine the velocity and position in the launch inertial frame at the moment of power-off in the powered deceleration phase. Step S2: Based on the position of the launch inertial frame at the moment the power deceleration phase is shut down, and the position of the target point in the launch frame, determine the absolute longitude difference between the position of the launch inertial frame at the moment the power deceleration phase is shut down and the target point. Step S3: Determine the remaining flight time of the passive segment based on the absolute longitude difference and the position of the target point in the launch inertial frame; Step S4: Determine the required speed and program angle calculation value for closed-circuit guidance based on the remaining flight time of the passive phase, the absolute longitude difference, and the speed in the launch inertial frame at the moment of shutdown of the powered deceleration phase.

[0024] An improved closed-loop guidance method for reusable rockets, according to embodiments of the present invention, employs a calculation method based on predicting the shutdown point state. This method can calculate the required velocity for closed-loop guidance, constraining the position and velocity direction of the passive phase endpoint after rocket shutdown. It is applicable to the powered deceleration phase of reusable rockets. This guidance method calculates the required velocity by predicting the shutdown point state, reducing the rate of change of the programmed angle during powered deceleration, optimizing control performance, and improving control stability. It can solve the closed-loop guidance algorithm parameters under non-minimum energy trajectory constraints, enabling the engineering application of the closed-loop guidance algorithm in the powered deceleration phase during rocket first-stage recovery. Furthermore, this guidance method is not limited to the reentry deceleration phase of reusable rockets; it is applicable in scenarios requiring control of the passive phase endpoint after shutdown, such as high-altitude recovery phases or rocket ascent phases. It achieves high guidance accuracy, has a simple algorithm structure, high computational efficiency, and is easy to run on different hardware.

[0025] Example 1: According to an embodiment of the present invention, in step S1, the velocity and position in the launch inertial frame at the moment of shutdown of the powered deceleration phase are determined based on the initial value of the remaining flight time of the active phase, the real-time gravitational acceleration, and the initial value of the program angle of the powered deceleration phase. This includes: determining the remaining flight time of the active phase based on the initial value of the remaining flight time of the active phase and the number of operation cycles; setting the differential equation to be solved based on the real-time gravitational acceleration and the initial value of the program angle of the powered deceleration phase; and determining the velocity and position in the launch inertial frame at the moment of shutdown of the powered deceleration phase based on the remaining flight time of the active phase and the differential equation to be solved.

[0026] According to an embodiment of the present invention, when calculating the remaining flight time of the active phase, the remaining flight time of the active phase can be determined according to the following formula (1). : (1) According to an embodiment of the present invention, the flight phase in which the rocket engine provides thrust is the active phase. In formula (1), This represents the initial value of the remaining flight time in the active phase, which is a pre-determined and fixed parameter (i.e., a parameter that is calculated in advance through ground simulation). t represents the computation cycle duration, is the binding value, and n represents the current computation cycle number, starting from 1. In another example, the initial value of the remaining flight time of the active segment can also be trained on a neural network based on a large amount of offline sample data, and then predicted by the neural network based on multiple parameters of the active segment, thereby further improving the accuracy of the initial value of the remaining flight time of the active segment and the accuracy of subsequent calculations.

[0027] According to an embodiment of the present invention, after obtaining the remaining flight time of the active phase, a differential equation can be set to solve for the velocity in the launch inertial frame at the moment of shutdown during the powered deceleration phase. , , ) and location ( The differential equation to be solved can be set according to the following formula (2): (2) According to an embodiment of the present invention, the reusable rocket's re-entry phase during its return, where the engine reignites to decelerate and adjust its trajectory, is termed the powered deceleration phase. In formula (2), 'a' represents the average apparent acceleration during the powered deceleration phase, which is a bound value. , as well as It is the real-time gravitational acceleration in the x, y, and z axes calculated based on the flight status. , This is the program angle for the power deceleration phase. In the first calculation cycle, it is the set value, and in subsequent calculation cycles, it is the closed-circuit guidance solution value from the previous cycle. for Regarding the differential value of time, for Regarding the differential value of time, for Regarding the differential value of time, for Regarding the differential value of time, for Regarding the differential value of time, for Differential value with respect to time.

[0028] According to an embodiment of the present invention, the velocity and position in the launch inertial frame at the moment of engine shutdown during the active phase are determined based on the remaining flight time of the active phase and the differential equation to be solved. The launch inertial frame, at the instant of rocket liftoff, has its three axes perfectly aligned with the launch frame at that time and is fixedly pointing towards a distant star, independent of Earth's rotation. The lower limit of integration is set to 0, and the upper limit of integration is set to the remaining flight time of the active phase. And solve the above differential equation using Euler's algorithm. , , )and( This allows us to obtain the velocity in the launch inertial frame at the moment of shutdown during the power deceleration phase. ) and location ( ).

[0029] In this way, the remaining flight time of the active segment for each frame can be calculated by using a pre-loaded value decreasing the calculation cycle. Furthermore, a neural network can be fitted based on the offline simulation results of a large sample size, and the remaining flight time can be calculated by using the input flight state, thereby improving the accuracy of the prediction.

[0030] Example 2: According to an embodiment of the present invention, in step S2, determining the absolute longitude difference between the position in the launch inertial frame at the moment of shutdown of the power deceleration phase and the target point, based on the position of the target point in the launch frame and the position of the launch point in the launch frame, includes: obtaining the geocentric radius vector of the origin of the launch inertial frame and the transformation matrix from the launch inertial frame to the launch frame; determining the geocentric radius vector of the predicted shutdown point based on the position in the launch inertial frame at the moment of shutdown of the power deceleration phase, the geocentric radius vector of the origin of the launch inertial frame, and the transformation matrix from the launch inertial frame to the launch frame; determining the geocentric radius vector of the target point based on the geocentric radius vector of the origin of the launch inertial frame and the position of the target point in the launch frame; determining the geocentric latitude of the predicted shutdown point and the geocentric latitude of the target point based on the geocentric radius vector of the predicted shutdown point and the geocentric radius vector of the target point; and determining the absolute longitude difference between the position in the launch inertial frame and the target point at the moment of shutdown of the power deceleration phase based on the geocentric latitude of the predicted shutdown point, the geocentric latitude of the target point, the geographical latitude of the launch point, and the geocentric radius vector of the predicted shutdown point.

[0031] According to an embodiment of the present invention, the geocentric radius vector of the origin of the launch inertial frame and the transformation matrix from the launch inertial frame to the launch frame can be obtained. For example, the time elapsed after rocket liftoff is t, and the Earth's rotational angular velocity is... The transformation matrix from the launching inertial frame to the launching frame. ,For example, This can be used to convert between the launch inertial frame and the launch frame. Since the origin of the launch inertial frame is the launch point at the moment of rocket liftoff, the geocentric radius at the launch point is the same as the geocentric radius at the origin of the launch inertial frame. .

[0032] According to an embodiment of the present invention, based on the position in the launch inertial frame at the moment of shutdown during the power deceleration phase, the geocentric radius vector of the origin of the launch inertial frame, and the transformation matrix from the launch inertial frame to the launch frame obtained above, the geocentric radius vector of the predicted shutdown point can be calculated. In the example, the geocentric radius vector of the predicted shutdown point can be determined according to formula (3). , (3) According to an embodiment of the present invention, the launch system is a rectangular coordinate system fixed at the launch point (e.g., the position where the rocket takes off) and rotating with the Earth's rotation. The X-axis usually points in the direction of the launch azimuth (i.e., the target direction facing the rocket after takeoff), the Y-axis usually points vertically to the zenith, and the Z-axis usually forms a right-handed coordinate system with the X and Y axes, pointing to the side.

[0033] According to an embodiment of the present invention, the geocentric radius of the target point can be determined based on the geocentric radius of the origin of the launching inertial frame and the position of the target point in the launching frame. In the example, the geocentric radius of the target point can be determined according to formula (4). , (4) According to an embodiment of the present invention, in formula (4), This represents the three-axis components of the target point in the launch system, and is a bound value.

[0034] According to an embodiment of the present invention, the geocentric latitude of the predicted shutdown point and the geocentric latitude of the target point are determined based on the geocentric radius vector of the predicted shutdown point and the geocentric radius vector of the target point. In the example, the geocentric latitude of the predicted shutdown point can be determined according to formula (5). and the geocentric latitude of the target point , (5) According to an embodiment of the present invention, in formula (5), This represents the vector of Earth's rotational angular velocity in the launch inertial frame. express The modulus of the two vectors, the " " indicates dot product.

[0035] Furthermore, based on the geocentric latitude of the predicted shutdown point, the geocentric latitude of the target point, the geographical latitude of the launch point, and the geocentric radius vector of the target point, the absolute longitude difference between the position in the launch inertial frame at the moment of shutdown during the powered deceleration phase and the target point can be determined. In the example, the absolute longitude difference between the position in the launch inertial frame at the moment of shutdown during the powered deceleration phase and the target point can be determined according to formula (6). and , (6) According to an embodiment of the present invention, in formula (6), Indicates the geographical latitude of the launch point. for The z-axis components after normalization for The normalized x-axis components, for The normalized x-axis components, for The z-axis components after normalization for The normalized x-axis components, for The z-axis component after normalization.

[0036] Example 3: According to an embodiment of the present invention, in step S3, determining the remaining flight time of the passive segment based on the absolute longitude difference and the position of the target point under the launch system includes: setting a constructor based on the absolute longitude difference and the position of the target point under the launch system; converting the constructor into an iterative function in iterative form; iteratively solving the iterative function, and obtaining the remaining flight time of the passive segment when the iterative function satisfies a preset convergence condition.

[0037] According to an embodiment of the present invention, a constructor is set based on the absolute longitude difference and the position of the target point in the launch system, including: determining the range angle based on the absolute longitude difference; determining the trajectory inclination angle based on the position of the target point in the launch system and the transformation matrix from the launch system to the launch inertial system at the end of the passive phase; determining the passive phase orbital semi-circle based on the range angle and the trajectory inclination angle; determining the predicted shutdown point and the apogee angle of the target point based on the passive phase orbital semi-circle, the trajectory inclination angle, and the range angle; determining the orbital parameters based on the apogee angle and the passive phase orbital semi-circle; and setting the constructor based on the orbital parameters.

[0038] According to an embodiment of the present invention, the range angle can be calculated based on the geocentric latitude of the predicted shutdown point, the geocentric latitude of the target point, the absolute longitude difference, and other parameters determined above. In the example, the range angle can be determined according to formula (7). , (7) According to an embodiment of the present invention, the range angle It can describe the geocentric angle between the predicted shutdown point and the target point, and can describe the angle that the velocity vector should satisfy when the rocket reaches the target point (e.g., requiring the rocket to land vertically). In formula (7), This represents the result of the remaining flight time iteration for the i-th iteration, with the initial value being the binding value.

[0039] Furthermore, the trajectory inclination angle, i.e., the angle between the trajectory and the horizontal plane, can be determined based on the position of the target point in the launch frame and the transformation matrix from the launch frame to the launch inertial frame at the end of the passive phase. In the example, the trajectory inclination angle can be determined according to formulas (8) and (9). : (8) (9) According to an embodiment of the present invention, in formula (8), This represents the transformation matrix from the launch frame to the launch inertial frame at the end of the passive phase, and can be used to realize the transformation from the launch frame to the launch inertial frame. , , This represents the three-axis components of the desired velocity vector at the end of the passive phase in the launch frame. It is a bound value used to constrain the velocity direction during rocket landing; for example, to achieve a vertical soft landing, it can be set... =0、 =0、 =0. Furthermore, the above range angle and trajectory inclination angle are calculated without aerodynamic influence. In order to further improve the calculation accuracy, aerodynamic influence compensation terms can be added through simulation and other means to obtain the range angle and trajectory inclination angle considering aerodynamic influence.

[0040] According to an embodiment of the present invention, the passive section trajectory semi-circular diameter can be determined based on the range angle and trajectory inclination angle determined above. In the example, the passive section trajectory semi-circular diameter can be determined according to formula (10). That is, the vertical distance from the orbital focus (e.g., the Earth's center of mass) to the orbital curve.

[0041] (10) According to an embodiment of the present invention, the predicted shutdown point and the apogee angle of the target point can be determined based on the passive section trajectory semi-circle, ballistic inclination angle, and range angle determined above. In the example, the predicted shutdown point apogee angle can be determined according to formulas (11), (12), and (13). Angle of far point from the target point .

[0042] (11) (12) (13).

[0043] Furthermore, based on the apogee angle and the semi-circular diameter of the passive section track determined above, multiple track parameters can be determined. In the example, the eccentricity e and the apogee angle of the predicted shutdown point can be determined according to formula (14). The deflection angle of the target point at its farthest point and semi-major axis.

[0044] (14) According to an embodiment of the present invention, the remaining flight time can be calculated based on Kepler's third law according to formula (15). The remaining flight time iteration result in this round of iteration is as follows: The value of .

[0045] (15) Where GM is the Earth's gravitational constant.

[0046] According to an embodiment of the present invention, an iterative function can be constructed based on Newton's iteration method according to formula (16), and the initial iterative function can be determined by numerical differentiation. (16) Furthermore, by taking a step size of 0.01, the constructor can be determined according to formula (17). (17) According to an embodiment of the present invention, when When the time is less than 0.1, the iteration converges, and the remaining flight time of the passive segment is output. The segment where the rocket glides without power is the passive segment. If If the value is greater than or equal to 0.1, the remaining flight time for the next iteration is calculated using formula (17), substituted into formula (7), and a new iteration is performed.

[0047] In this way, Newton's iteration can be used to calculate nonlinear equations, which enhances the robustness and speed of the solution process. It can also be combined with the two-body problem to solve some of the equations analytically, expanding the application scenarios of closed-loop guidance algorithms under non-minimum energy orbit conditions, and enabling closed-loop guidance algorithms to meet the guidance requirements of reusable rocket propulsion deceleration phases.

[0048] Example 4: According to an embodiment of the present invention, in step S4, determining the required velocity for closed-circuit guidance and the calculated program angle value based on the remaining flight time of the passive phase, the absolute longitude difference, and the velocity in the launch inertial frame at the moment of shutdown of the powered deceleration phase includes: determining the required velocity for closed-circuit guidance based on the remaining flight time of the passive phase and the absolute longitude difference; and determining the calculated program angle value based on the required velocity for closed-circuit guidance and the velocity in the launch inertial frame at the moment of shutdown of the powered deceleration phase.

[0049] According to an embodiment of the present invention, determining the required velocity for closed-circuit guidance based on the remaining flight time and absolute longitude difference of the passive phase includes: determining the required velocity modulus and the predicted trajectory inclination in the shutdown state based on the passive phase orbital semi-circularity, orbital parameters, and trajectory inclination; determining the angle between the orbital plane and the meridional plane based on the absolute longitude difference, the remaining flight time of the passive phase, and the range angle; and determining the required velocity for closed-circuit guidance based on the angle between the orbital plane and the meridional plane, the required velocity modulus, and the predicted trajectory inclination in the shutdown state.

[0050] According to an embodiment of the present invention, the required velocity modulus can be determined based on the passive section track semi-circularity, track parameters, and trajectory inclination, for example, formula (18). , (18) Furthermore, the predicted trajectory angle in the shutdown state can be calculated based on the required velocity modulus determined above. For example, the predicted trajectory angle in the shutdown state can be determined according to formula (19). It can describe the angle between the rocket's velocity vector and the local horizontal plane at the predicted engine shutdown time.

[0051] (19) According to an embodiment of the present invention, the angle between the orbital plane and the meridian plane can be determined based on the absolute longitude difference, the remaining flight time of the passive segment, and the range angle determined above. In the example, the angle between the orbital plane and the meridian plane can be determined according to formula (20). , (20) According to an embodiment of the present invention, the angle between the orbital plane and the meridian plane can describe the angle between the plane containing the unpowered flight trajectory of the rocket after shutdown and the longitude plane passing through the Earth's center and the target point.

[0052] Furthermore, the required velocity for closed-circuit guidance can be determined based on the angle between the orbital plane and the meridional plane, the required velocity modulus, and the predicted trajectory inclination angle in the shutdown state. In the example, the required velocity for closed-circuit guidance can be determined according to formula (21). ,in, , , These are the three-axis components of the velocity required for closed-circuit guidance. The calculation methods for multiple parameters in formula (21) are shown in formula (22). (twenty one) (twenty two) According to an embodiment of the present invention, through the above calculations, the required velocity for closed-circuit guidance can be decomposed into horizontal velocity components within the orbital plane. Vertical velocity component and the horizontal velocity component outside the orbital plane The complex constraints in three-dimensional space are decomposed into three mutually orthogonal axes with clear physical meaning.

[0053] According to an embodiment of the present invention, the program angle calculation value can be determined based on the required speed for closed-circuit guidance and the velocity in the launch inertial frame at the moment of shutdown during the powered deceleration phase. In the example, the program angle calculation value can be determined according to formula (23). (twenty three) In the example, after the program angle calculation is completed, the program angle instruction can be output to the control system, and the calculated value of the program angle can be stored for prediction of the shutdown state of the power deceleration stage in the next operation cycle.

[0054] The method described in this application (B) allows for prediction of the shutdown state during the powered deceleration phase using a constant program angle assumption. It also considers different application scenarios or flight trajectories, improving the adaptability of the closed-loop guidance algorithm to the powered deceleration phase. Furthermore, it can be further improved based on different program angle assumptions, extending its application to other stages of rocket flight. Moreover, this guidance method has a simple algorithm structure, high computational efficiency, and is easy to run on different hardware.

[0055] Example 5: Figure 2 A flowchart of an improved closed-circuit guidance method for reusable rockets according to an embodiment of the present invention is shown as an example.

[0056] like Figure 2 As shown, based on the initial value of the remaining flight time in the active phase, the real-time gravitational acceleration, and the initial value of the program angle in the powered deceleration phase, the velocity and position in the launch inertial frame at the moment of shutdown in the powered deceleration phase are determined. That is, according to the basic principle of the closed-loop guidance algorithm, the rocket velocity at the moment of shutdown is equal to the required velocity for closed-loop guidance (i.e., the terminal velocity constraint must be satisfied). Therefore, according to optimal control theory, the optimal program angle constant for the deceleration phase is loaded. Under the condition of known program angle and remaining flight time in the active phase, the rocket velocity and position at the moment of shutdown in the powered deceleration phase are calculated by integration. Specifically, the program angle used for the first shutdown state prediction is a pre-set value; subsequent shutdown state predictions use the program angle output from the previous closed-loop guidance cycle. The remaining flight time in the active phase used for the first shutdown state prediction is a pre-set value; subsequent remaining flight time in the active phase used for the current shutdown state prediction is obtained by subtracting the algorithm's running cycle from the remaining flight time in the previous active phase. Then, based on the position in the launch inertial frame at the moment of shutdown during powered deceleration, and the position of the target point in the launch frame, the absolute longitude difference between the position in the launch inertial frame and the target point at the moment of shutdown during powered deceleration is determined. This determines the remaining flight time in the passive phase, and subsequently, the required velocity and programmed angle for closed-circuit guidance are determined. In other words, after calculating the predicted position at shutdown using the previous section, the predicted state is used to perform closed-circuit guidance calculations to obtain the required velocity.

[0057] Since the required velocity cannot be directly calculated using analytical formulas when the flight trajectory does not meet the minimum energy orbit constraint, an iterative method is used to calculate the orbital semi-circle and the remaining flight time of the passive segment based on the current position, target position, and target velocity direction. This determines the elliptical trajectory of the rocket after shutdown, and the required velocity is calculated based on the orbital parameters. Furthermore, the gain velocity is obtained by subtracting the required velocity from the predicted shutdown point velocity, and the program angle is calculated using the gain velocity.

[0058] Example 6: Figure 3 An exemplary block diagram of an improved closed-loop guidance system for a reusable rocket according to an embodiment of the present invention is shown, the system comprising: The velocity and position module determines the velocity and position in the launch inertial frame at the moment of shutdown during the powered deceleration phase, based on the initial value of the remaining flight time during the active phase, the real-time gravitational acceleration, and the initial value of the programmed angle during the powered deceleration phase. The absolute longitude difference module determines the absolute longitude difference between the position in the launch inertial frame at the moment the power deceleration phase is shut down and the position of the target point in the launch frame. The remaining flight time module determines the remaining flight time of the passive segment based on the absolute longitude difference and the position of the target point under the launch system; The required speed and program angle module determines the required speed and program angle values ​​for closed-circuit guidance based on the remaining flight time of the passive phase, the absolute longitude difference, and the velocity in the launch inertial frame at the moment of shutdown in the powered deceleration phase.

[0059] According to one embodiment of the present invention, an improved closed-circuit guidance device suitable for reusable rockets is provided, comprising: a processor; a memory for storing processor-executable instructions; wherein the processor is configured to invoke the instructions stored in the memory to execute the improved closed-circuit guidance method suitable for reusable rockets.

[0060] According to one embodiment of the present invention, a computer-readable storage medium is provided having computer program instructions stored thereon, which, when executed by a processor, implement the improved closed-circuit guidance method applicable to reusable rockets.

[0061] This invention can be a method, apparatus, system, and / or computer program product. The computer program product may include a computer-readable storage medium having computer-readable program instructions loaded thereon for performing various aspects of the invention.

[0062] Those skilled in the art should understand that the embodiments of the present invention described above and shown in the accompanying drawings are merely examples and do not limit the present invention. The objectives of the present invention have been fully and effectively achieved. The functions and structural principles of the present invention have been demonstrated and explained in the embodiments, and any variations or modifications may be made to the implementation of the present invention without departing from the stated principles.

[0063] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some or all of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the embodiments of the present invention.

Claims

1. An improved closed loop guidance method suitable for reusable rockets, characterized in that, include: Based on the initial value of the remaining flight time in the active phase, the real-time gravitational acceleration, and the initial value of the programmed angle in the powered deceleration phase, determine the velocity and position in the launch inertial frame at the moment of shutdown in the powered deceleration phase; Based on the position of the target point in the launch inertial frame at the moment of shutdown of the power deceleration phase, and the position of the target point in the launch frame, determine the absolute longitude difference between the position of the target point in the launch inertial frame at the moment of shutdown of the power deceleration phase. The remaining flight time of the passive segment is determined based on the absolute longitude difference and the position of the target point under the launch system. Based on the remaining flight time of the passive phase, the absolute longitude difference, and the velocity in the launch inertial frame at the moment of shutdown in the powered deceleration phase, determine the required velocity and program angle solution value for closed-loop guidance; Based on the position of the target point in the launch inertial frame at the moment of shutdown during the powered deceleration phase, and the position of the target point in the launch frame, determine the absolute longitude difference between the position of the target point in the launch inertial frame at the moment of shutdown during the powered deceleration phase, including: Obtain the geocentric radius vector of the origin of the launch inertial frame and the transformation matrix from the launch inertial frame to the launch frame; Based on the position in the launch inertial frame at the moment of shutdown during the power deceleration phase, the geocentric radius of the origin of the launch inertial frame, and the transformation matrix from the launch inertial frame to the launch frame, the geocentric radius of the predicted shutdown point is determined. The geocentric radius of the target point is determined based on the geocentric radius of the origin of the launch inertial frame and the position of the target point in the launch frame. Based on the geocentric radius vector of the predicted shutdown point and the geocentric radius vector of the target point, determine the geocentric latitude of the predicted shutdown point and the geocentric latitude of the target point; Based on the geocentric latitude of the predicted shutdown point, the geocentric latitude of the target point, the geographical latitude of the launch point, and the geocentric radius vector of the predicted shutdown point, determine the absolute longitude difference between the launch inertial frame position and the target point at the moment of shutdown during the powered deceleration phase.

2. The modified closed path guidance method suitable for reusable rockets as claimed in claim 1, wherein, Based on the initial value of the remaining flight time in the active phase, the real-time gravitational acceleration, and the initial value of the programmed angle in the powered deceleration phase, determine the velocity and position in the launch inertial frame at the moment of power-off in the powered deceleration phase, including: The remaining flight time of the active phase is determined based on the initial value of the remaining flight time of the active phase and the number of calculation cycles. Based on the real-time gravitational acceleration and the initial value of the program angle during the dynamic deceleration phase, the differential equation to be solved is set. Based on the remaining flight time of the active phase and the differential equations to be solved, determine the velocity and position in the launch inertial frame at the moment of shutdown during the powered deceleration phase.

3. The improved closed-circuit guidance method for reusable rockets according to claim 1, characterized in that, Based on the absolute longitude difference and the target point's position within the launch system, the remaining flight time for the passive segment is determined, including: Based on the absolute longitude difference and the position of the target point in the launching system, set the constructor; Transform the constructor into an iterative function of iterative form; The iterative function is solved iteratively, and the remaining flight time of the passive segment is obtained if the iterative function meets the preset convergence condition.

4. The improved closed-circuit guidance method for reusable rockets according to claim 3, characterized in that, Based on the absolute longitude difference and the target point's position in the launch system, a constructor is set, including: Determine the range angle based on the absolute longitude difference; The trajectory inclination angle is determined based on the position of the target point in the launch frame and the transformation matrix from the launch frame to the launch inertial frame at the end of the passive phase. Determine the passive section trajectory semi-bore based on the range angle and the trajectory inclination angle; Based on the passive section trajectory semi-circle, ballistic inclination angle, and range angle, determine the predicted shutdown point and the apogee angle of the target point; The track parameters are determined based on the apogee angle and the semi-circular diameter of the passive track segment; Based on the orbital parameters, set the constructor.

5. The improved closed-circuit guidance method for reusable rockets according to claim 4, characterized in that, Based on the remaining flight time of the passive phase, the absolute longitude difference, and the velocity in the launch inertial frame at the moment of shutdown in the powered deceleration phase, the required velocity and program angle calculation values ​​for closed-loop guidance are determined, including: The required speed for closed-loop guidance is determined based on the remaining flight time and absolute longitude difference of the passive segment. The program angle calculation value is determined based on the required speed for closed-circuit guidance and the speed in the launch inertial frame at the moment of shutdown during the power deceleration phase.

6. The improved closed-circuit guidance method for reusable rockets according to claim 5, characterized in that, Based on the remaining flight time and absolute longitude difference of the passive segment, determine the required velocity for closed-circuit guidance, including: Based on the passive section track semi-circle, track parameters, and ballistic inclination, determine the required velocity modulus and predict the ballistic inclination in the shutdown state; The angle between the orbital plane and the meridian plane is determined based on the absolute longitude difference, the remaining flight time of the passive segment, and the range angle. The required velocity for closed-circuit guidance is determined based on the angle between the orbital plane and the meridional plane, the required velocity modulus, and the predicted trajectory inclination angle in the shutdown state.

7. An improved closed-circuit guidance system suitable for reusable rockets, characterized in that, include: The velocity and position module determines the velocity and position in the launch inertial frame at the moment of shutdown during the powered deceleration phase, based on the initial value of the remaining flight time during the active phase, the real-time gravitational acceleration, and the initial value of the programmed angle during the powered deceleration phase. The absolute longitude difference module determines the absolute longitude difference between the position in the launch inertial frame at the moment the power deceleration phase is shut down and the position of the target point in the launch frame. The remaining flight time module determines the remaining flight time of the passive segment based on the absolute longitude difference and the position of the target point under the launch system; The required speed and program angle module determines the required speed and program angle values ​​for closed-loop guidance based on the remaining flight time of the passive phase, the absolute longitude difference, and the velocity in the launch inertial frame at the moment of shutdown in the powered deceleration phase. Based on the position of the target point in the launch inertial frame at the moment of shutdown during the powered deceleration phase, and the position of the target point in the launch frame, determine the absolute longitude difference between the position of the target point in the launch inertial frame at the moment of shutdown during the powered deceleration phase, including: Obtain the geocentric radius vector of the origin of the launch inertial frame and the transformation matrix from the launch inertial frame to the launch frame; Based on the position in the launch inertial frame at the moment of shutdown during the power deceleration phase, the geocentric radius of the origin of the launch inertial frame, and the transformation matrix from the launch inertial frame to the launch frame, the geocentric radius of the predicted shutdown point is determined. The geocentric radius of the target point is determined based on the geocentric radius of the origin of the launch inertial frame and the position of the target point in the launch frame. Based on the geocentric radius vector of the predicted shutdown point and the geocentric radius vector of the target point, determine the geocentric latitude of the predicted shutdown point and the geocentric latitude of the target point; Based on the geocentric latitude of the predicted shutdown point, the geocentric latitude of the target point, the geographical latitude of the launch point, and the geocentric radius vector of the predicted shutdown point, determine the absolute longitude difference between the launch inertial frame position and the target point at the moment of shutdown during the powered deceleration phase.

8. An improved closed-circuit guidance device suitable for reusable rockets, characterized in that, include: processor; A memory for storing processor-executable instructions; wherein the processor is configured to invoke the instructions stored in the memory to perform the method as described in any one of claims 1-6.

9. A computer-readable storage medium, characterized in that, It stores computer program instructions that, when executed by a processor, implement the method of any one of claims 1-6.

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