Carrier rocket reuse transverse movement control instruction compensation method based on position information feedback

Abstract

The invention relates to a carrier rocket reuse control instruction compensation method based on current position output feedback. The carrier rocket reuse control instruction compensation method comprises the following steps: (1) giving definition of a launching coordinate system and representation of a thrust vector in the coordinate system; (2) generating a required control instruction according to current and expected speed and position information; (3) filtering the position information by using a mean filtering algorithm, and generating a correction amount of the control instruction in the step (2); (4) calculating the required thrust according to the control instructions in the step (2) and the step (3) and the correction amount thereof; and (5) obtaining a corresponding thrust instruction direction according to the thrust vector direction definition in the step (1). According to the method, the corresponding control instruction correction is generated according to the position deviation calculated by the flight software at the current moment, the position information is filtered by using the mean filtering algorithm, the correction of the control instruction is obtained, and the compensation effect of the method is more remarkable especially under the condition that large constant interference exists.

Description

TECHNICAL FIELD

[0001] The present application relates to a kind of high-precision landing point control technology of reusable launch site of carrier rocket, especially suitable for the control object with the nonlinear, uncertainty and strong coupling etc. Characteristic carrier rocket reuse control instruction compensation method. BACKGROUND

[0002] Low-cost access to space is the dream of mankind for a long time, and the recycling and reuse of launch vehicle and spacecraft is an important way to achieve this goal. This is not only the core demand of commercial space development, but also the key support for manned space and deep space exploration scientific tasks. Through the successful recycling and reuse of Falcon 9 sub-level, SpaceX has verified the feasibility of this technology in engineering practice. However, China still faces many challenges in rocket landing control technology: not only need to solve the online trajectory planning problem under multi-stage and strong constraint conditions, overcome the stability control problem caused by the strong coupling between guidance and attitude control, but also must realize the high-precision landing control of carrier rocket.

[0003] Currently, to achieve landing control, pseudo-spectral method, convex optimization, polynomial and other online calculation guidance methods are usually used to control the rocket to fly along the predetermined trajectory to the target point. Although this method is widely used in trajectory tracking control, the position control accuracy will be significantly affected when the system is disturbed, especially by constant disturbance, resulting in static deviation, and the greater the disturbance, the more obvious the deviation.

[0004] Therefore, the present application proposes a method for improving the position control accuracy of the aircraft based on position information feedback. The method generates a preliminary correction amount of control instruction based on the position deviation amount calculated by the flight software at the current time; then, the mean filter algorithm is applied to filter the position information to obtain the final accurate correction amount used for control. This scheme can effectively suppress the position error caused by disturbance, especially for constant disturbance, and shows significant disturbance adaptive control ability, that is, the greater the disturbance, the stronger the compensation ability. SUMMARY

[0005] The present application aims to improve the flight control of existing carrier rockets with nonlinear, uncertain and strongly coupled characteristics of the controlled object, and proposes a carrier rocket reuse control instruction compensation method based on current position output feedback. According to the current position information calculated by the flight software, the deviation amount of the current position and the position is calculated, and then the expression of the relationship between the corrected control instruction and the required thrust vector is obtained, which completes the improvement of the control instruction based on accelerometer output information feedback. Specifically, it includes the following steps:

[0006] (1) The definition of launch coordinate system and the representation of thrust vector in this coordinate system are given;

[0007] (2) generating the required control command according to the current and expected speed and position information;

[0008] (3) applying mean filter algorithm to filter the position information to generate the correction amount of the control command in step (2);

[0009] (4) calculating the required thrust according to the control command and its correction amount in step (2) and step (3);

[0010] (5) obtaining the corresponding thrust command direction according to the thrust vector direction definition in step (1).

[0011] Preferably, the launch coordinate system in step (1) is defined as follows: and the thrust vector in the launch coordinate system is defined as follows: is the center of mass of the aircraft, is the opposite direction of the direction of gravity at the launch moment of the aircraft, is the launch direction at the launch moment of the aircraft, and , form a right-handed orthogonal coordinate system, and the launch coordinate system moves with the earth rotation; the lateral movement control command referred to in the present application is the control command in the direction of the launch coordinate system and ; the projection of the thrust vector in is the angle , and counterclockwise rotation is positive; the projection of the thrust vector in is the angle , and outward is positive; the above-mentioned coordinate system, lateral movement control command, and thrust vector definition are not unique, and the user can define them according to the needs. Preferably, step (2) generates the required control command in the direction of the axis according to the current and expected speed and position information, which is:

[0012]

[0013] ;

[0014] the control command in the direction of the axis is:

[0015]

[0016] wherein: , , , , , ​​are the proportional and differential coefficients respectively; , , , are the and axis direction flight software calculated and expected position information, units are both m; , and , are the and axis direction flight software calculated and expected velocity information, units are both m / s; , are the and axis direction command acceleration, units are m / s 2 ; each parameter is represented in the launch coordinate system of the aircraft; , and , can be valued according to the user's needs, if landing at the take-off and landing point, they are all 0.

[0017] Preferably, the step (3) applies a mean filter algorithm to filter the position information, generating step (2) The correction amount of the control command in the axis direction is:

[0018] ;

[0019] The correction amount of the control command in the axis direction is:

[0020] ;

[0021] In the formula: , are the , axis direction command acceleration correction amount coefficients; , are the , axis direction command acceleration correction amount; , is the calculated value of the current position, which is:

[0022] ;

[0023] .

[0024] In the formula: N is the number of data points for averaging; in actual flight, due to the dynamic process of control and navigation errors in the current position calculation, the average value over a period of time is taken as the expected value calculated by the flight control software at the current moment during the stable speed phase.

[0025] Preferably, in step (4), the required thrust acceleration command is calculated based on the control commands and correction values ​​from steps (2) and (3):

[0026] ;

[0027] In the formula: subscript They represent , Parameters corresponding to the axial direction; The coefficient for compensation can be selected based on the characteristics of the engine, and is usually taken as 0.5; The gravitational acceleration of the spacecraft at the current moment is calculated using the following formula:

[0028] ;

[0029] in: , , , , , , , ; The gravitational constant of Earth has a value of 3.986005 × 10¹⁴ m. 3 / s 2 ; The Earth's J2 gravitational coefficient has a value of 0.00108263. The radius of the Earth's equator is 6,378,140 m. The azimuth angle for launch; The geodetic latitude of the launch point; , and The three-axis components of the position from the Earth's center to the launch point in the launch coordinate system; , and These are the 3-axis components of the spacecraft in the launch coordinate system.

[0030] From the revised thrust command The expression for the expected thrust magnitude is derived as follows:

[0031] ;

[0032] Preferably, step (5) is to derive the desired attitude angle of the thrust vector direction based on the coordinate system definition in step (1):

[0033] ;

[0034] in and The range of values ​​for are as follows:

[0035] ;

[0036] The desired thrust The thrust is sent to the power system to generate the corresponding thrust, and the direction of the thrust is... , The corresponding thrust vector is sent to the attitude control system.

[0037] Compared with the prior art, the present invention has the following advantages:

[0038] (1) High control accuracy: It can effectively improve the control accuracy under various disturbances such as thrust line deviation, center of mass deviation and aerodynamic force, and the stronger the disturbance, the more significant the improvement effect.

[0039] (2) Good compatibility: The new method is highly compatible with the original control scheme, and when the correction amount is zero, the new method automatically degenerates into the original method, which facilitates engineering integration and smooth transition;

[0040] (3) Strong implementation: The algorithm is simple and easy to implement in engineering, which is of great significance for achieving high-precision control of the landing point of the launch vehicle debris or the precise landing of the sub-stage. Attached Figure Description

[0041] Figure 1 This represents the thrust vector in the rocket body coordinate system. Detailed Implementation

[0042] The present invention will now be described in further detail with reference to the accompanying drawings and specific embodiments, including:

[0043] (1) The definition of the launch coordinate system and the representation of the thrust vector in this coordinate system are given as follows:

[0044] Launch coordinate system The thrust vector is defined in the launch coordinate system as follows: For the center of mass of the aircraft, The direction opposite to the direction of gravity at the moment of spacecraft launch. The firing direction at the moment of launch of the spacecraft. and , The axes form a right-handed orthogonal coordinate system, and the launch coordinate system moves with the Earth's rotation; the lateral control command referred to in this invention is the launch coordinate system. and Directional control commands; thrust vector exist The projection and The included angle is Counterclockwise rotation is positive; thrust vector exist The projection and The included angle between the planes is The outward direction is positive; the definitions of the above coordinate system, lateral control command, and thrust vector are not unique, and users can define them according to their needs.

[0045] (2) Generate based on current and desired speed and location information. The control commands required for the axial direction are:

[0046] ;

[0047] The direction control command is:

[0048] ;

[0049] in: , , , , , These are the proportional and differential coefficients, respectively; , , , They are respectively and The position information calculated and expected by the axial flight software, both in meters; , and , They are respectively and The calculated and expected speed information in the axial direction flight software, both in m / s; , for and Commanded acceleration in the axial direction, in m / s² 2 All parameters are represented in the spacecraft launch coordinate system. , and , The value can be selected according to the user's needs; if landing at the take-off and landing point, the value will be 0.

[0050] (3) Apply the mean filtering algorithm to filter the location information to generate step (2). The correction amount for the axial direction control command is:

[0051] ;

[0052] The correction amount for the axial direction control command is:

[0053] ;

[0054] In the formula: , for , The coefficient for the correction amount of the axial direction command acceleration; , for , Correction amount for axial direction command acceleration; , The calculated value for the current position is:

[0055] ;

[0056] .

[0057] In the formula: N is the number of data points for averaging; in actual flight, due to the dynamic process of control and navigation errors in the current position calculation, the average value over a period of time is taken as the expected value calculated by the flight control software at the current moment during the stable speed phase.

[0058] (4) Based on the control commands and their corrections from steps (2) and (3), calculate the required thrust acceleration command as follows:

[0059] ;

[0060] In the formula: subscript They represent , Parameters corresponding to the axial direction; The coefficient for compensation can be selected based on the characteristics of the engine, and is usually taken as 0.5; The gravitational acceleration of the spacecraft at the current moment is calculated using the following formula:

[0061] ;

[0062] in: , , , , , , , ; The gravitational constant of Earth has a value of 3.986005 × 10¹⁴ m. 3 / s 2 ; The Earth's J2 gravitational coefficient has a value of 0.00108263. The radius of the Earth's equator is 6,378,140 m. The azimuth angle for launch; The geodetic latitude of the launch point; , and The three-axis components of the position from the Earth's center to the launch point in the launch coordinate system; , and These are the 3-axis components of the spacecraft in the launch coordinate system.

[0063] From the revised thrust command The expression for the expected thrust magnitude is derived as follows:

[0064] ;

[0065] (5) Based on the definition of the thrust vector direction in step (1), the corresponding thrust command direction is obtained as follows:

[0066] ;

[0067] in and The range of values ​​for are as follows:

[0068] ;

[0069] The desired thrust The thrust is sent to the power system to generate the corresponding thrust, and the direction of the thrust is... , The corresponding thrust vector is sent to the attitude control system.

[0070] This invention generates corresponding control command corrections based on the position deviation calculated by the flight software at the current moment, and applies a mean filtering algorithm to filter the position information to obtain the control command correction, effectively suppressing position errors caused by interference. Especially in the presence of constant interference, the greater the interference, the more significant the compensation effect of this method.

[0071] The above description is only the best specific embodiment of the present invention, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in the present invention should be included within the protection scope of the present invention.

[0072] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A method for compensating reusable control commands for launch vehicles based on current position output feedback, characterized in that: Includes the following steps: (1) Give the definition of the launch coordinate system and the representation of the thrust vector in this coordinate system; Preferably, the launch coordinate system The thrust vector is defined in the launch coordinate system as follows: For the center of mass of the aircraft, The direction opposite to the direction of gravity at the moment of spacecraft launch. The firing direction at the moment of launch of the spacecraft. and , The axes form a right-handed orthogonal coordinate system, and the launch coordinate system moves with the Earth's rotation; the lateral control command referred to in this invention is the launch coordinate system. and Directional control commands; thrust vector exist The projection and The included angle is Counterclockwise rotation is positive; thrust vector exist The projection and The included angle between the planes is Outward is positive; the definitions of the above coordinate system, lateral control command, and thrust vector are not unique, and users can define them as needed. (2) Generate the required control commands based on the current and desired speed and position information; (3) Apply the mean filtering algorithm to filter the position information and generate the correction amount of the control command in step (2); (4) Calculate the required thrust based on the control commands and corrections from steps (2) and (3); (5) Based on the thrust vector direction definition in step (1), the corresponding thrust command direction is obtained.

2. The method for compensating reusable launch vehicle control commands based on current position output feedback according to claim 1, characterized in that... Step (2) Generate the required speed and location information based on the current and desired speed and location information. The control command for the axial direction is ； The direction control command is ； in: , , , , , These are the proportional and differential coefficients, respectively; , , , They are respectively and The position information calculated and expected by the axial flight software, both in meters; , and , They are respectively and The calculated and expected speed information in the axial direction flight software, both in m / s; , for and Commanded acceleration in the axial direction, in m / s² 2 All parameters are represented in the spacecraft launch coordinate system. , and , The value can be selected according to the user's needs; if landing at the take-off and landing point, the value will be 0.

3. The method for compensating reusable launch vehicle control commands based on current position output feedback according to claim 1, characterized in that... Step (3) applies the mean filtering algorithm to filter the location information to generate step (2). The correction amount for the axial direction control command is ； The correction amount for the axial direction control command is ； In the formula: , for , The coefficient for the correction amount of the axial direction command acceleration; , for , Correction amount for axial direction command acceleration; , The calculated value for the current position is... ； ； In the formula: N is the number of data points for which the average value is taken; In actual flight, due to the dynamic process of control and navigation errors in the current position calculation, the average value over a period of time is taken as the expected value calculated by the flight control software at the current moment during the stable speed phase.

4. The method for compensating reusable launch vehicle control commands based on current position output feedback according to claim 1, characterized in that... Step (4) involves calculating the required thrust acceleration command based on the control commands and corrections from steps (2) and (3). ； In the formula: subscript They represent , Parameters corresponding to the axial direction; The coefficient for compensation can be selected based on the characteristics of the engine, and is usually taken as 0.5; The gravitational acceleration of the spacecraft at the current moment is calculated using the following formula: ； in: , , , , , , , ; The gravitational constant of Earth has a value of 3.986005 × 10¹⁴ m. 3 / s 2 ; The Earth's J2 gravitational coefficient has a value of 0.00108263. The radius of the Earth's equator is 6,378,140 m. The azimuth angle for launch; The geodetic latitude of the launch point; , and The three-axis components of the position from the Earth's center to the launch point in the launch coordinate system; , and The three-axis components of the spacecraft in the launch coordinate system; derived from the corrected thrust command. The expression for the expected thrust magnitude is derived as follows: 。 5. The method for compensating reusable launch vehicle control commands based on current position output feedback according to claim 1, characterized in that... Step (5) is based on the coordinate system definition in step (1) to obtain the desired attitude angle of the thrust vector direction as follows: ； in and The range of values ​​for are as follows: ； The desired thrust The thrust is sent to the power system to generate the corresponding thrust, and the direction of the thrust is... , The corresponding thrust vector is sent to the attitude control system.

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