A rocket return guidance method, device, equipment and medium

Abstract

The present invention discloses a rocket return guidance method, device, equipment and medium, which determine the radius vector difference between the current position of the rocket and the position of the target point, as well as the current flight speed of the rocket; determine the target celestial acceleration of the rocket according to the radius vector difference, the current flight speed of the rocket and the non-integer power of time; determine the target thrust of the rocket engine according to the target celestial acceleration and the predicted mass of the rocket; use a polynomial guidance algorithm to solve the target north acceleration and the target east acceleration, and determine the program attitude angle of the rocket in combination with the target thrust, so that the attitude control system of the rocket can return the rocket according to the program attitude angle. The present application determines the celestial acceleration by a non-integer power formula. Compared with the traditional method of determining the acceleration by a quadratic polynomial, the present application can constrain the thrust ratio within a certain range, so that the thrust adjustment range is small, which improves the reliability of the engine operation and reduces the difficulty of engine research and development.

Description

Technical Field

[0001] The present invention relates to the field of rocket technology, and in particular to a rocket return guidance method, device, equipment and medium. Background Art

[0002] The flight environment of reusable liquid rockets changes highly dynamically. During the landing process, the guidance law gives the thrust vector direction, thrust size and engine working time, so that the rocket can suppress the influence of disturbance deviation and accurately land vertically at the predetermined position. This vertical landing guidance problem that relies on vector thrust back thrust and considers position, speed and attitude constraints is the powered descent guidance problem.

[0003] In related technologies, polynomial guidance methods are usually used to achieve vertical soft landing of aircraft. However, the conventional polynomial guidance requires a relatively large engine thrust adjustment range. The larger the thrust adjustment range, the deeper the engine thrust will be required to be throttled, which usually leads to problems such as reduced combustion efficiency and increased structural vibration. In severe cases, unstable combustion, deteriorated heat transfer, low-frequency coupled vibration and other faults may occur, affecting the safety of rocket flight and reducing the reliability of engine operation. Therefore, the polynomial guidance method has the defect of being unable to constrain the upper and lower limits of the thrust size, which increases the difficulty of engine development. Summary of the invention

[0004] The embodiments of the present application provide a rocket return guidance method, device, equipment and medium, which solves the technical problem in the prior art that the use of traditional polynomial guidance leads to a relatively large engine thrust adjustment range, thereby reducing the engine operating reliability, and achieves the technical effect of improving the engine operating reliability.

[0005] In a first aspect, the present application provides a rocket return guidance method, the method comprising:

[0006] Determine the radius difference between the current position of the rocket and the position of the target point, as well as the current flight speed of the rocket;

[0007] Determine the target celestial acceleration of the rocket based on the radius vector difference, the current flight speed of the rocket and the non-integer power of time;

[0008] Determine the target thrust of the rocket's engine based on the target celestial acceleration and the predicted mass of the rocket;

[0009] A polynomial guidance algorithm is used to solve the target north acceleration and target east acceleration. The rocket's programmed attitude angle is determined based on the target north acceleration, target east acceleration and target thrust, so that the rocket's attitude control system can guide the rocket back according to the programmed attitude angle.

[0010] Furthermore, the radius vector difference between the current position of the rocket and the position of the target point is determined, including:

[0011] Get the first geocentric vector radius of the rocket's current location and the second geocentric vector radius of the rocket's target point;

[0012] Project the first geocentric radius vector and the second geocentric radius vector to the north celestial east coordinate system of the target point to determine the radius vector difference between the current position of the rocket and the position of the target point.

[0013] Furthermore, the current flight speed of the rocket is determined, including:

[0014] Determine the rocket's current speed in the North Celestial East coordinate system at the target point.

[0015] Furthermore, the target celestial acceleration is:

[0016]

[0017] Among them, a1(t) is the celestial acceleration, C0 and C1 are constant vectors to be determined, and t is time.

[0018] Furthermore, a polynomial guidance algorithm is used to solve the target north acceleration and the target east acceleration, and the program attitude angle of the rocket is determined according to the target north acceleration, the target east acceleration and the target thrust, including:

[0019] The polynomial guidance algorithm is used to determine the target north acceleration and the target east acceleration of the rocket in the north celestial east coordinate system of the target point;

[0020] The programmed attitude angle of the rocket is determined based on the target thrust, target north acceleration, and target east acceleration.

[0021] Further, the polynomial guidance algorithm is:

[0022] a2(t)=C2+C3t+C4t 2

[0023] Among them, a2(t) is the celestial acceleration, C2, C3, C4 are constant vectors to be determined, and t is time.

[0024] In a second aspect, the present application provides a rocket return guidance device, the device comprising:

[0025] The first determination module is used to determine the radius difference between the current position of the rocket and the position of the target point, as well as the current flight speed of the rocket;

[0026] The second determination module is used to determine the target celestial acceleration of the rocket according to the radius vector difference, the current flight speed of the rocket and the non-integer power of time;

[0027] The third determination module is used to determine the target thrust of the rocket's engine according to the target celestial acceleration and the predicted mass of the rocket;

[0028] The fourth determination module is used to use a polynomial guidance algorithm to solve the target north acceleration and the target east acceleration, and determine the rocket's programmed attitude angle based on the target north acceleration, the target east acceleration and the target thrust, so that the rocket's attitude control system can guide the rocket back according to the programmed attitude angle.

[0029] Furthermore, the first determination module includes:

[0030] The acquisition submodule is used to obtain the first geocentric vector radius of the current position of the rocket and the second geocentric vector radius of the target point of the rocket;

[0031] The projection submodule is used to project the first geocentric radius vector and the second geocentric radius vector into the North Celestial East coordinate system of the target point to determine the radius vector difference between the current position of the rocket and the position of the target point.

[0032] In a third aspect, the present application provides an electronic device, including:

[0033] processor;

[0034] a memory for storing processor-executable instructions;

[0035] Wherein, the processor is configured to execute to implement a rocket return guidance method.

[0036] In a fourth aspect, the present application provides a non-temporary computer-readable storage medium, which, when the instructions in the storage medium are executed by a processor of an electronic device, enables the electronic device to implement a rocket return guidance method.

[0037] One or more technical solutions provided in the embodiments of the present application have at least the following technical effects or advantages:

[0038] This application first determines the radius vector difference between the current position of the rocket and the position of the target point, as well as the current flight speed of the rocket, and then determines the target celestial acceleration of the rocket based on the radius vector difference, the current flight speed of the rocket and the non-integer power of time, and then determines the target thrust and attitude angle of the rocket to achieve return guidance. Therefore, this application determines the celestial acceleration by a non-integer power formula. Compared with the traditional method of determining acceleration by a quadratic polynomial, this embodiment can constrain the thrust ratio within a certain range, making the thrust adjustment range smaller, thereby improving the reliability of the engine and reducing the difficulty of engine research and development. BRIEF DESCRIPTION OF THE DRAWINGS

[0039] In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the following briefly introduces the drawings required for use in the description of the embodiments. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0040] Figure 1 A schematic diagram of a flow chart of a rocket return guidance method provided in this application;

[0041] Figure 2 A schematic diagram of a thrust curve determined according to a non-integer power formula provided in this application;

[0042] Figure 3 is a schematic diagram of a thrust curve determined according to a quadratic polynomial;

[0043] Figure 4 A schematic diagram of the structure of a rocket return guidance device provided in this application;

[0044] Figure 5 A schematic diagram of the structure of an electronic device provided in this application. DETAILED DESCRIPTION

[0045] The embodiment of the present application provides a rocket return guidance method, which solves the technical problem in the prior art that the use of traditional polynomial guidance results in a relatively large engine thrust adjustment range, thereby reducing the engine operating reliability.

[0046] The technical solution of the embodiment of the present application is to solve the above technical problems, and the overall idea is as follows:

[0047] A rocket return guidance method comprises: determining the radius vector difference between the current position of the rocket and the position of the target point, and the current flight speed of the rocket; determining the target celestial acceleration of the rocket according to the radius vector difference, the current flight speed of the rocket and the non-integer power of time; determining the target thrust of the rocket engine according to the target celestial acceleration and the predicted mass of the rocket; using a polynomial guidance algorithm to solve the target north acceleration and the target east acceleration, and determining the program attitude angle of the rocket according to the target north acceleration, the target east acceleration and the target thrust, so that the attitude control system of the rocket can perform return guidance on the rocket according to the program attitude angle.

[0048] The present application determines the celestial acceleration through the non-integer power formula of the radius vector difference, the current flight speed of the rocket and the time. Compared with the traditional quadratic polynomial to determine the acceleration, the present embodiment can constrain the thrust ratio within a certain range, so that the thrust adjustment range is smaller, thereby improving the engine working reliability and reducing the difficulty of engine research and development.

[0049] In order to better understand the above technical solution, the above technical solution will be described in detail below in conjunction with the accompanying drawings and specific implementation methods.

[0050] First of all, the term "and / or" in this article is only a description of the association relationship of the associated objects, indicating that there can be three relationships. For example, A and / or B can mean: A exists alone, A and B exist at the same time, and B exists alone. In addition, the character " / " in this article generally indicates that the associated objects before and after are in an "or" relationship.

[0051] This embodiment provides Figure 1 A rocket return guidance method is shown, the method comprising steps S11 to S14.

[0052] Step S11, determining the radius vector difference between the current position of the rocket and the position of the target point, as well as the current flight speed of the rocket.

[0053] Specifically, by obtaining the first geocentric vector radius of the rocket at its current position and the second geocentric vector radius of the rocket's target point; projecting the first geocentric vector radius and the second geocentric vector radius to the North Celestial East coordinate system of the target point, the radius vector difference between the rocket's current position and the target point's position is determined.

[0054] For example, the first geocentric radius is set to R btd0 , the second geocentric radius is set to R btdf , then the formula for the radius vector difference is as follows:

[0055] ΔR btd =R btd0 -R btdf

[0056] To determine the current flight speed of the rocket is essentially to obtain the actual speed of the rocket at its current position, project the actual speed into the North Celestial East coordinate system of the target point, and then obtain the current flight speed of the rocket in the North Celestial East coordinate system of the target point.

[0057] Step S12, determining the target celestial acceleration of the rocket based on the radius vector difference, the current flight speed of the rocket and the non-integer power of time.

[0058] The non-integer power formula of the target's azimuthal acceleration with respect to time is shown in formula (1).

[0059]

[0060] Among them, a1(t) is the celestial acceleration, C0 and C1 are constant vectors to be determined, and t is time.

[0061] According to the above formula (1), if the current time is taken as time 0, that is, t is 0, then the celestial acceleration in formula (1) is C0. Therefore, to calculate the celestial acceleration at the current time, it is only necessary to determine C0. The calculation method of C0 is as follows:

[0062] By integrating formula (1), we can obtain the speed formula (2) shown below.

[0063]

[0064] By integrating formula (2), we can obtain the position formula (3) shown below.

[0065]

[0066] The rocket return guidance method provided in this embodiment is mainly used in the movement process between the rocket and the target point. Therefore, the initial position r0, initial velocity v0 of the rocket at the current position, and the target position r0 when the rocket lands f and target speed v f are all known, so, combined with the initial position r0, initial velocity v0, target position r f and target speed v f According to formula (2) and (3), the formula for C0 can be determined as follows:

[0067]

[0068] Therefore, the celestial acceleration can be determined as:

[0069]

[0070] Among them, y represents the direction of the sky, a y is the celestial acceleration, v btd0y Indicates the current celestial velocity, v btdfy Indicates the target speed in the sky, ΔR btdy represents the celestial radial vector difference, t go The remaining flight time.

[0071] In this embodiment, the celestial acceleration is determined by a non-integer power formula as shown in formula (1), which can be seen from Figure 2 As shown ( Figure 2 In the figure, the horizontal axis is time and the vertical axis is the ratio of the actual thrust P to the full thrust P0). Compared with the traditional quadratic polynomial to determine the acceleration (see Figure 3 As shown in FIG. 1 , where the horizontal axis is time and the vertical axis is the ratio of the actual thrust P to the full thrust P0), this embodiment can constrain the thrust ratio within a certain range, while Figure 3In the example, both the maximum value and the minimum value must be limited, which makes it difficult to determine the upper and lower limits of the thrust range, making the thrust adjustment range larger. When the thrust adjustment range is larger, the engine thrust will be required to be deeply throttled, which usually leads to problems such as reduced combustion efficiency and aggravated structural vibration. In severe cases, unstable combustion, deteriorated heat transfer, low-frequency coupled vibration and other faults may occur, affecting the flight safety of the rocket, reducing the reliability of the engine, and increasing the difficulty of engine research and development. Therefore, this embodiment determines the celestial acceleration through a non-integer power formula, which can greatly reduce the difficulty of engine research and development.

[0072] Step S13, determining the target thrust F of the rocket's engine according to the target celestial acceleration and the predicted mass of the rocket.

[0073] F=ma ybtd

[0074] a ybtd =a y -g ybtd

[0075] Among them, g ybtd is the projection of the gravitational acceleration in the celestial direction in the North Celestial East coordinate system, and m is the predicted mass.

[0076] Step S14, using a polynomial guidance algorithm to solve the target north acceleration and the target east acceleration, and determining the rocket's programmed attitude angle based on the target north acceleration, the target east acceleration and the target thrust, so that the rocket's attitude control system can guide the rocket back according to the programmed attitude angle.

[0077] A polynomial guidance algorithm is used to solve the target north acceleration and the target east acceleration, and the program attitude angle of the rocket is determined according to the target north acceleration, the target east acceleration and the target thrust, which specifically includes steps S21 and S22.

[0078] Step S21, using a polynomial guidance algorithm to determine the target north acceleration and target east acceleration of the rocket in the north celestial east coordinate system of the target point.

[0079] Preset quadratic polynomials, including:

[0080] a2(t)=C2+C3t+C4t 2 (6)

[0081] Among them, a2(t) is the celestial acceleration, C2, C3, C4 are constant vectors to be determined, and t is time.

[0082] According to the above formula (6), if the current time is 0, that is, t is 0, then a2(t) in formula (6) is C2, and a2(t) represents the northward acceleration or the eastward acceleration. Therefore, to calculate the northward acceleration or the eastward acceleration at the current time, it is only necessary to determine C2. The calculation method of C2 is as follows:

[0083] By integrating formula (6), we can obtain the speed formula (7) shown below.

[0084]

[0085] By integrating formula (7), we can obtain the position formula (8) shown below.

[0086]

[0087] The rocket return guidance method provided in this embodiment is mainly used in the movement process between the rocket and the target point. Therefore, the initial position r0, initial velocity v0 of the rocket at the current position, and the target position r0 when the rocket lands f 、Target speed v f and the target acceleration a f (When landing, the acceleration is 0) are all known, so combined with the initial position r0, initial velocity v0, the target position r f and target speed v f 、Target acceleration a f According to formula (7) and (8), the formula for C2 can be determined as follows:

[0088]

[0089] Therefore, the target north acceleration and the target east acceleration are shown in formulas (10) and (11), respectively.

[0090]

[0091]

[0092] Among them, x represents north direction, a x represents the northward acceleration, v btd0x Indicates the current northward speed, v btdfx Indicates the northward target speed, ΔR btdx represents the north radial difference, t go is the remaining flight time. z represents eastward, a z is the eastward acceleration, v btd0z represents the current eastward speed, v btdfz represents the eastward target speed, ΔR btdz represents the eastward radial difference, t goThe remaining flight time.

[0093] Step S22, determining the programmed attitude angle of the rocket based on the target thrust, the target north acceleration and the target east acceleration.

[0094] a xbtd =a x -g xbtd

[0095] a zbtd =a z -g zbtd

[0096] [a xdh a ydh a zdh ] T =L bdt2dh [a xbtd a ybtd a zbtd ] T

[0097]

[0098]

[0099] Among them, g xbtd , g zbtd is the projection of gravity acceleration in the north and east directions in the Beitiandong coordinate system, a xdh 、a ydh 、a zdh is the projection of the acceleration required for guidance in the north, sky and east directions in the north-sky-east coordinate system, L bdt2dh is the transformation matrix from the North Celestial East coordinate system to the navigation coordinate system; ψ dh It is the pitch and yaw program attitude angle in the navigation system.

[0100] In summary, in this embodiment, the celestial acceleration is determined by a non-integer power formula as shown in formula (1), which can be seen from Figure 2 Compared with the traditional quadratic polynomial to determine acceleration (see Figure 3 As shown in FIG. 1 , this embodiment can constrain the thrust ratio within a certain range, and Figure 3In the example, both the maximum value and the minimum value must be limited, which makes it difficult to determine the upper and lower limits of the thrust range, making the thrust adjustment range larger. When the thrust adjustment range is larger, the engine thrust will be required to be deeply throttled, which usually leads to problems such as reduced combustion efficiency and aggravated structural vibration. In severe cases, unstable combustion, deteriorated heat transfer, low-frequency coupled vibration and other faults may occur, affecting the flight safety of the rocket, reducing the reliability of the engine, and increasing the difficulty of engine research and development. Therefore, this embodiment determines the celestial acceleration through a non-integer power formula, which can greatly reduce the difficulty of engine research and development.

[0101] Based on the same inventive concept, this embodiment provides Figure 4 A rocket return guidance device, comprising:

[0102] The first determination module 41 is used to determine the radius vector difference between the current position of the rocket and the position of the target point, and the current flight speed of the rocket;

[0103] The second determination module 42 is used to determine the target celestial acceleration of the rocket according to the radius vector difference, the current flight speed of the rocket and the non-integer power of time; the non-integer power formula of the target celestial acceleration with respect to time includes:

[0104]

[0105] Among them, a1(t) is the celestial acceleration, C0 and C1 are constant vectors to be determined, and t is time.

[0106] A third determination module 43 is used to determine a target thrust of the rocket's engine according to the target celestial acceleration and the predicted mass of the rocket;

[0107] The fourth determination module 44 is used to use a polynomial guidance algorithm to solve the target north acceleration and the target east acceleration, and determine the rocket's program attitude angle according to the target north acceleration, the target east acceleration and the target thrust, so that the rocket's attitude control system can return the rocket according to the program attitude angle. The preset quadratic polynomial includes:

[0108] a2(t)=C2+C3t+C4t 2

[0109] Among them, a2(t) is the celestial acceleration, C2, C3, C4 are constant vectors to be determined, and t is time.

[0110] The first determining module 41 includes:

[0111] The acquisition submodule is used to obtain the first geocentric vector radius of the current position of the rocket and the second geocentric vector radius of the target point of the rocket;

[0112] The projection submodule is used to project the first geocentric radius vector and the second geocentric radius vector into the North Celestial East coordinate system of the target point to determine the radius vector difference between the current position of the rocket and the position of the target point.

[0113] The first determining module 41 includes:

[0114] The first determination submodule is used to determine the current flight speed of the rocket in the north celestial east coordinate system of the target point.

[0115] The fourth determination module includes:

[0116] The acceleration determination submodule is used to determine the target north acceleration and target east acceleration of the rocket in the north celestial east coordinate system at the target point by using a polynomial guidance algorithm;

[0117] The attitude angle determination submodule is used to determine the programmed attitude angle of the rocket based on the target thrust, target north acceleration and target east acceleration.

[0118] Specifically, by obtaining the first geocentric vector radius of the rocket at its current position and the second geocentric vector radius of the rocket's target point; projecting the first geocentric vector radius and the second geocentric vector radius to the North Celestial East coordinate system of the target point, the radius vector difference between the rocket's current position and the target point's position is determined.

[0119] For example, the first geocentric radius is set to R btd0 , the second geocentric radius is set to R btdf , then the formula for the radius vector difference is as follows:

[0120] ΔR btd =R btd0 -R btdf

[0121] To determine the current flight speed of the rocket is essentially to obtain the actual speed of the rocket at its current position, project the actual speed into the North Celestial East coordinate system of the target point, and then obtain the current flight speed of the rocket in the North Celestial East coordinate system of the target point.

[0122] The non-integer power formula of the preset celestial acceleration is shown in formula (12).

[0123]

[0124] Among them, a1(t) is the celestial acceleration, C0 and C1 are constant vectors to be determined, and t is time.

[0125] According to the above formula (12), the current time is taken as time 0, that is, t is 0, then the celestial acceleration in formula (12) is C0. Therefore, to calculate the celestial acceleration at the current time, it is only necessary to determine C0. The calculation method of C0 is as follows:

[0126] By integrating formula (12), we can obtain the speed formula (13) shown below.

[0127]

[0128] By integrating formula (13), we can obtain the position formula (14) shown below.

[0129]

[0130] The rocket return guidance method provided in this embodiment is mainly used in the movement process between the rocket and the target point. Therefore, the initial position r0, initial velocity v0 of the rocket at the current position, and the target position r0 when the rocket lands f and target speed v f are all known, so, combined with the initial position r0, initial velocity v0, the target position r f and target speed v f According to formula (13) (14), the formula for C0 can be determined as follows:

[0131]

[0132] Therefore, the celestial acceleration can be determined as:

[0133]

[0134] Among them, y represents the direction of the sky, a y is the celestial acceleration, v btd0y Indicates the current celestial velocity, v btdfy Indicates the target speed in the sky, ΔR btdy represents the celestial radial vector difference, t go The remaining flight time.

[0135] In this embodiment, the celestial acceleration is determined by a non-integer power formula as shown in formula (12), which can be seen from Figure 2 Compared with the traditional quadratic polynomial to determine acceleration (see Figure 3 As shown in FIG. 1 , this embodiment can constrain the thrust ratio within a certain range, and Figure 3In the example, both the maximum value and the minimum value must be limited, which makes it difficult to determine the upper and lower limits of the thrust range, making the thrust adjustment range larger. When the thrust adjustment range is larger, the engine thrust will be required to be deeply throttled, which usually leads to problems such as reduced combustion efficiency and aggravated structural vibration. In severe cases, unstable combustion, deteriorated heat transfer, low-frequency coupled vibration and other faults may occur, affecting the flight safety of the rocket, reducing the reliability of the engine, and increasing the difficulty of engine research and development. Therefore, this embodiment determines the celestial acceleration through a non-integer power formula, which can greatly reduce the difficulty of engine research and development.

[0136] The polynomial guidance algorithm includes:

[0137] a2(t)=C2+C3t+C4t 2 (17)

[0138] Among them, a2(t) is the celestial acceleration, C2, C3, C4 are constant vectors to be determined, and t is time.

[0139] According to the above formula (17), if the current time is taken as time 0, that is, t is 0, then a2(t) in formula (17) is C2, and a2(t) represents the northward acceleration or the eastward acceleration. Therefore, to calculate the northward acceleration or the eastward acceleration at the current time, it is only necessary to determine C2. The calculation method of C2 is as follows:

[0140] By integrating formula (17), we can obtain the speed formula (18) shown below.

[0141]

[0142] By integrating formula (18), we can obtain the position formula (19) shown below.

[0143]

[0144] The rocket return guidance method provided in this embodiment is mainly used in the movement process between the rocket and the target point. Therefore, the initial position r0, initial velocity v0 of the rocket at the current position, and the target position r0 when the rocket lands f 、Target speed v f and the target acceleration a f (When landing, the acceleration is 0) are all known, so combined with the initial position r0, initial velocity v0, the target position r f and target speed v f 、Target acceleration a f According to formula (18) (19), the formula for C2 can be determined as follows:

[0145]

[0146] Therefore, the target north acceleration and the target east acceleration are shown in formulas (21) and (22), respectively.

[0147]

[0148]

[0149] Among them, x represents north direction, a x represents the northward acceleration, v btd0x Indicates the current northward speed, v btdfx Indicates the northward target speed, ΔR btdx Represents the north radial difference, t go is the remaining flight time. z represents eastward, a z is the eastward acceleration, v btd0z represents the current eastward speed, v btdfz represents the eastward target speed, ΔR btdz represents the eastward radial difference, t go The remaining flight time.

[0150] In summary, in this embodiment, the celestial acceleration is determined by a non-integer power formula as shown in formula (12), which can be seen from Figure 2 Compared with the traditional quadratic polynomial to determine acceleration (see Figure 3 As shown in FIG. 1 , this embodiment can constrain the thrust ratio within a certain range, and Figure 3 In the example, both the maximum value and the minimum value must be limited, which makes it difficult to determine the upper and lower limits of the thrust range, making the thrust adjustment range larger. When the thrust adjustment range is larger, the engine thrust will be required to be deeply throttled, which usually leads to problems such as reduced combustion efficiency and aggravated structural vibration. In severe cases, unstable combustion, deteriorated heat transfer, low-frequency coupled vibration and other faults may occur, affecting the flight safety of the rocket, reducing the reliability of the engine, and increasing the difficulty of engine research and development. Therefore, this embodiment determines the celestial acceleration through a non-integer power formula, which can greatly reduce the difficulty of engine research and development.

[0151] Based on the same inventive concept, this embodiment provides Figure 5 An electronic device as shown includes:

[0152] Processor 51;

[0153] A memory 52 for storing instructions executable by the processor 51;

[0154] The processor 51 is configured to execute to implement a rocket return guidance method.

[0155] Based on the same inventive concept, this embodiment provides a non-temporary computer-readable storage medium. When the instructions in the storage medium are executed by the processor 51 of the electronic device, the electronic device can execute a rocket return guidance method.

[0156] Since the electronic device introduced in this embodiment is an electronic device used to implement the information processing method in the embodiment of the present application, based on the information processing method introduced in the embodiment of the present application, a person skilled in the art can understand the specific implementation of the electronic device of the present embodiment and its various variations, so how the electronic device implements the method in the embodiment of the present application is not described in detail here. As long as a person skilled in the art implements the electronic device used by the information processing method in the embodiment of the present application, it belongs to the scope of protection of this application.

[0157] Those skilled in the art will appreciate that embodiments of the present invention may be provided as methods, systems, or computer program products. Therefore, the present invention may take the form of a complete hardware embodiment, a complete software embodiment, or an embodiment combining software and hardware. Moreover, the present invention may take the form of a computer program product implemented on one or more computer-usable storage media (including but not limited to disk storage, CD-ROM, optical storage, etc.) containing computer-usable program code.

[0158] The present invention is described with reference to flowcharts and / or block diagrams of methods, devices (systems), and computer program products according to embodiments of the present invention. It should be understood that each process and / or block in the flowchart and / or block diagram, as well as the combination of processes and / or blocks in the flowchart and / or block diagram, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, a special-purpose computer, an embedded processor, or other programmable data processing device to produce a machine, so that the instructions executed by the processor of the computer or other programmable data processing device generate instructions for implementing the processes in the flowchart and / or block diagram. Figure 1 A process or multiple processes and / or boxes Figure 1 A device that provides the functions specified in a block or multiple blocks.

[0159] These computer program instructions may also be stored in a computer-readable memory capable of directing a computer or other programmable data processing device to operate in a specific manner, so that the instructions stored in the computer-readable memory produce an article of manufacture comprising an instruction device, which implements the process Figure 1 A process or multiple processes and / or boxes Figure 1 A function specified in one or more boxes.

[0160] These computer program instructions can also be loaded onto a computer or other programmable data processing device so that a series of operating steps are executed on the computer or other programmable device to produce a computer-implemented process, thereby providing instructions for implementing the process. Figure 1 A process or multiple processes and / or boxes Figure 1 The steps for the functions specified in one or more boxes.

[0161] Although the preferred embodiments of the present invention have been described, those skilled in the art may make other changes and modifications to these embodiments once they have learned the basic creative concept. Therefore, the appended claims are intended to be interpreted as including the preferred embodiments and all changes and modifications that fall within the scope of the present invention.

[0162] Obviously, those skilled in the art can make various changes and modifications to the present invention without departing from the spirit and scope of the present invention. Thus, if these modifications and variations of the present invention fall within the scope of the claims of the present invention and their equivalents, the present invention is also intended to include these modifications and variations.

Claims

1. A rocket return guidance method, characterized in that: The method comprises: Determine the radius vector difference between the current position of the rocket and the position of the target point, and the current flight speed of the rocket; Determining the target celestial acceleration of the rocket according to the radius vector difference, the current flight speed of the rocket and the non-integer power of time; Determining a target thrust of an engine of the rocket according to the target celestial acceleration and the predicted mass of the rocket; A polynomial guidance algorithm is used to solve the target north acceleration and the target east acceleration, and the programmed attitude angle of the rocket is determined according to the target north acceleration, the target east acceleration and the target thrust, so that the attitude control system of the rocket can return the rocket according to the programmed attitude angle.

2. The method according to claim 1, characterized in that Determining the radius vector difference between the current position of the rocket and the position of the target point includes: Obtaining a first geocentric vector radius of the current position of the rocket and a second geocentric vector radius of the target point of the rocket; The first geocentric radius vector and the second geocentric radius vector are projected into the North Celestial East coordinate system of the target point to determine the radius vector difference between the current position of the rocket and the position of the target point.

3. The method according to claim 1, characterized in that Determining the current flight speed of the rocket includes: Determine the current flight speed of the rocket at the target point in the north celestial east coordinate system.

4. The method according to claim 1, characterized in that The target celestial acceleration is: Among them, the is the celestial acceleration, C0 and C1 are constant vectors to be determined, For time.

5. The method according to claim 1, characterized in that The method adopts a polynomial guidance algorithm to solve the target north acceleration and the target east acceleration, and determines the program attitude angle of the rocket according to the target north acceleration, the target east acceleration and the target thrust, including: Determine the target north acceleration and the target east acceleration of the rocket in the north celestial east coordinate system of the target point by using the polynomial guidance algorithm; The programmed attitude angle of the rocket is determined based on the target thrust, the target northward acceleration and the target eastward acceleration.

6. The method according to claim 5, characterized in that The polynomial guidance algorithm is: Among them, the is the celestial acceleration, C2, C3, C4 are constant vectors to be determined, For time.

7. A rocket return guidance device, characterized in that: The device comprises: The first determination module is used to determine the radius difference between the current position of the rocket and the position of the target point, and the current flight speed of the rocket; A second determination module is used to determine the target celestial acceleration of the rocket according to the radius vector difference, the current flight speed of the rocket and the non-integer power of time; A third determination module is used to determine a target thrust of the engine of the rocket according to the target celestial acceleration and the predicted mass of the rocket; The fourth determination module is used to use a polynomial guidance algorithm to solve the target north acceleration and the target east acceleration, and determine the program attitude angle of the rocket based on the target north acceleration, the target east acceleration and the target thrust, so that the attitude control system of the rocket can return the rocket based on the program attitude angle.

8. The device according to claim 7, characterized in that The first determining module includes: An acquisition submodule, used to acquire a first geocentric vector radius of the current position of the rocket and a second geocentric vector radius of the target point of the rocket; The projection submodule is used to project the first geocentric radius vector and the second geocentric radius vector into the North Celestial East coordinate system of the target point to determine the radius vector difference between the current position of the rocket and the position of the target point.

9. An electronic device, characterized in that: include: processor; a memory for storing instructions executable by the processor; Wherein, the processor is configured to execute to implement a rocket return guidance method as described in any one of claims 1 to 6.

10. A non-temporary computer-readable storage medium, when the instructions in the storage medium are executed by a processor of an electronic device, enables the electronic device to implement a rocket return guidance method as described in any one of claims 1 to 6.

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