Carrier rocket reuse control instruction compensation method based on accelerometer output feedback

Abstract

The invention relates to a carrier rocket reuse control instruction compensation method based on accelerometer output feedback. The carrier rocket reuse control instruction compensation method comprises the following steps that (1) definition of a rocket body coordinate system and representation of a thrust vector in the coordinate system are given; (2) calculating measurement output information of the accelerometer in a rocket body coordinate system; (3) calculating the deviation between the measured output of the accelerometer and an instruction value, and filtering the deviation by using a mean filtering algorithm to provide more accurate input for control compensation; (4) according to the deviation value of the expected thrust and the actual thrust, the correction value of the acceleration instruction is calculated; and (5) obtaining a corresponding thrust instruction direction according to the thrust vector direction definition in the step (1). According to the measurement output of the aircraft accelerometer, the control algorithm is improved based on accelerometer output information feedback, and the method has great significance in improving flight control of a controlled object with the characteristics of nonlinearity, uncertainty, strong coupling and the like.

Description

Technical Field

[0001] This invention relates to the field of reusable control technology for high-precision precise landing of launch vehicles, and is particularly applicable to reusable control command compensation methods for launch vehicles with nonlinear, uncertain, and strongly coupled characteristics of the controlled object. Background Technology

[0002] With the rapid development of aerospace technology, the demand for high-precision automatic control of large, complex systems with strong uncertainties is becoming increasingly urgent. Existing online trajectory planning and control methods, such as pseudospectral methods, convex optimization, and polynomial methods, can achieve trajectory planning and control of aircraft, but their control effectiveness often declines when dealing with highly nonlinear or highly uncertain controlled objects, sometimes even leading to mission failure. Take the hovering mission as an example: this mission requires the thrust generated by the engine to balance the aircraft's gravity in real time. However, due to the inherent uncertainties in the propulsion system, the actual thrust often deviates from the expected value, and this deviation increases over time. In extreme cases, this could lead to the aircraft crashing or prolonged ascent, resulting in mission failure.

[0003] To address the aforementioned issues, the field of system identification and control has been dedicated to researching diagnostic methods for model uncertainties and exploring how to utilize diagnostic information to reduce the impact of uncertainties on system performance, achieving a series of advancements. Against this backdrop, this invention proposes a control command compensation method based on accelerometer output feedback. This method calculates the deviation between the actual acceleration and the commanded acceleration based on the actual measurement output of the aircraft's accelerometer, and then compensates the control command in real time according to the magnitude of this deviation.

[0004] The proposed method is applicable to control scenarios where the controlled object has significant nonlinearity or strong uncertainty. It can effectively improve the control performance of the aircraft, and its effect is more significant when the actual acceleration deviates from the commanded acceleration. In particular, it has important value in the design of flight control systems that deal with complex situations such as model nonlinearity, uncertainty and strong coupling. Summary of the Invention

[0005] The purpose of this invention is to improve the flight control of existing launch vehicles when the controlled object has nonlinear, uncertain, and strongly coupled characteristics. It proposes a reusable control command compensation method for launch vehicles based on accelerometer output feedback. Based on the measured output of the spacecraft's accelerometer, the deviation between the actual acceleration and the commanded acceleration is calculated. An expression for the relationship between the corrected control command and the required thrust vector is given, thus completing the improvement of the control command based on accelerometer output feedback. The specific steps include:

[0006] (1) Give the definition of the arrow body coordinate system and the representation of the thrust vector in this coordinate system;

[0007] (2) Calculate the measurement output information of the accelerometer in the rocket body coordinate system;

[0008] (3) Calculate the deviation between the accelerometer measurement output and the command value, and apply the mean filtering algorithm to filter the deviation to provide a more accurate input for control compensation;

[0009] (4) Calculate the correction amount of the acceleration command based on the deviation between the expected thrust and the actual thrust;

[0010] (5) Based on the thrust vector direction definition in step (1), the corresponding thrust command direction is obtained.

[0011] Preferably, the arrow body coordinate system in step (1) The thrust vector is defined in the rocket body coordinate system as follows: For the center of mass of the aircraft, The direction of the spacecraft's axis is opposite to the direction of gravity at the moment of launch. The launch time is in the opposite direction to the direction the spacecraft is heading. and , The axes form a right-handed orthogonal coordinate system; the 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 coordinate system and the representation of the thrust vector are not unique and can be defined by the user according to their needs.

[0012] Preferably, step (2) calculates the measurement output information of the accelerometer in the rocket body coordinate system as follows:

[0013] ;

[0014] in: , , , , and The coefficient of the first term of the accelerometer, , , They are respectively axis, shaft and The output of the shaft accelerometer; , , They are respectively axis, shaft and The zero-order term coefficient of the axial accelerometer; The calculation cycle for the flight control system software; , and After error compensation axis, shaft and Accelerometer output information in the axial direction.

[0015] Preferably, zero-term error refers to the fact that when the accelerometer has zero input, the accelerometer will have a certain non-zero volt output voltage due to manufacturing errors; first-term error refers to the fact that during the manufacturing process, even if the manufacturer has done precision machining, the sensitive axis of the accelerometer will usually deviate from the mounting axis by a small angle. These errors are usually obtained through single-machine calibration; the purpose of error compensation is to improve the navigation accuracy during on-orbit operation, and it is also a necessary task for all models at present.

[0016] Preferably, step (3) uses the average value of the accelerometer measurement output from step (2) to perform the following filtering calculation:

[0017] ;

[0018] In the formula: N is the number of data points for which the average value is taken; subscript They represent , or axis; The accelerometer measurement output value after error compensation in step (2); The desired acceleration value can be obtained by differentiating the user-planned flight trajectory.

[0019] Preferably, in step (4), based on the deviation between the desired thrust and the actual thrust, the following acceleration command is given. Correction calculation:

[0020] ;

[0021] In the formula: This represents the current mass of the aircraft.

[0022] Preferably, step (5) is to derive the modified thrust command based on step (4). for:

[0023] ;

[0024] In the formula: 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:

[0025] ;

[0026] 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.

[0027] From the revised thrust command The expression for the expected thrust magnitude is derived as follows:

[0028] ;

[0029] Based on the coordinate system definition in step (1), the desired attitude angle of the thrust vector direction is obtained as follows:

[0030] ;

[0031] in and The range of values ​​for are as follows:

[0032] ;

[0033] 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.

[0034] The advantages of this invention compared to the prior art are as follows:

[0035] (1) It can improve the control performance of the controlled object under nonlinear, uncertain and strongly coupled characteristics;

[0036] (2) By controlling the acceleration, the system can "sense" and "prepare" to deal with the upcoming dynamic demands before the speed changes significantly, which is especially suitable for situations with a high rate of change;

[0037] (3) It achieves great compatibility with the original method. When the correction amount is set to zero, the new method degenerates into the original method.

[0038] (4) The algorithm is simple and easy to implement in engineering. It is of great significance for large, complex and uncertain systems, especially for the flight control of reusable launch vehicles. Attached Figure Description

[0039] Figure 1 This represents the thrust vector in the rocket body coordinate system.

[0040] Figure 2 This is a schematic diagram of the calculation process of the present invention. Detailed Implementation

[0041] The present invention will now be described in further detail with reference to the accompanying drawings and specific embodiments, including:

[0042] (1) Give the definition of the rocket body coordinate system and the representation of the thrust vector in this coordinate system.

[0043] The coordinate system and thrust vector representation are not unique and can be defined by the user according to their needs. The arrow body coordinate system of this invention... The thrust vector is defined in the rocket body coordinate system as follows: For the center of mass of the aircraft, The direction of the spacecraft's axis is towards the direction of gravity at the moment of launch. The launch time is in the opposite direction to the direction the spacecraft is heading. and , The axes form a right-handed orthogonal coordinate system; the 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 , facing outwards is positive.

[0044] (2) Calculate the measurement output information of the accelerometer in the rocket body coordinate system, which is expressed as:

[0045] ;

[0046] in: , , , , and The coefficient of the first term of the accelerometer, , , They are respectively axis, shaft and The output of the shaft accelerometer; , , They are respectively axis, shaft and The zero-order term coefficient of the axial accelerometer; The calculation cycle for the flight control system software; , and After error compensation axis, shaft and Accelerometer output information in the axial direction.

[0047] Zero-term error refers to the non-zero volt output voltage of the accelerometer when the input is zero, due to manufacturing errors. First-term error refers to the fact that even with precision machining by the manufacturer, the sensitive axis of the accelerometer will usually deviate from the mounting axis by a small angle during the manufacturing process. These errors are usually obtained through single-unit calibration. The purpose of error compensation is to improve the navigation accuracy during on-orbit operation, and it is a necessary task for all models.

[0048] (3) Calculate the deviation between the accelerometer measurement output and the command value, and apply the mean filtering algorithm to filter the deviation.

[0049] ;

[0050] In the formula: N is the number of data points for which the average value is taken; see the table below. They represent , or ; The accelerometer measurement output value after error compensation; This represents the desired acceleration value.

[0051] (4) Calculate the correction amount of the acceleration command based on the deviation between the expected thrust and the actual thrust. for

[0052] ;

[0053] In the formula: This represents the current mass of the aircraft.

[0054] (5) Based on the definition of the thrust vector direction in step (1), the corresponding thrust command direction is obtained.

[0055] The corrected thrust command is obtained based on step (4). for:

[0056] ;

[0057] In the formula: 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:

[0058] ;

[0059] 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.

[0060] From the revised thrust command The expression for the expected thrust magnitude is derived as follows:

[0061] ;

[0062] Based on the coordinate system definition in step (1), the desired attitude angle of the thrust vector direction is obtained as follows:

[0063] ;

[0064] in and The range of values ​​for are as follows:

[0065] ;

[0066] 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.

[0067] This invention calculates the deviation between the actual acceleration and the commanded acceleration based on the measurement output of the aircraft's accelerometer, defines the attitude angle of the desired thrust in the rocket body coordinate system, and gives an expression for the relationship between the corrected control command and the required thrust vector. It completes the improvement of the control algorithm based on the feedback of accelerometer output information, and has great significance for improving the flight control of controlled objects with nonlinear, uncertain, and strongly coupled characteristics.

[0068] 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.

[0069] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A launch vehicle reusability control command compensation method based on accelerometer output feedback, characterized by: The method comprises the following steps: (1) defining the coordinate system of the arrow body and the representation of the thrust vector in the coordinate system; The body coordinate system of the arrow body is preferably defined as follows: The definitions of the thrust vector in the body coordinate system are as follows: is the mass center of the aircraft, is the direction of the aircraft axis, and is in the opposite direction of the gravity direction at the moment of launching, is in the opposite direction of the direction in which the aircraft is launched at the moment of launching, and , the axis forms a right-hand orthogonal coordinate system; the projection of the thrust vector on is the angle between the projection of the thrust vector on , and the clockwise rotation is positive; the projection of the thrust vector on is the angle between the projection of the thrust vector on , and the outward rotation is positive; the representation of the coordinate system and the thrust vector is not unique, and can be defined by the user as needed; (2) calculating the measurement output information of the accelerometer in the arrow body coordinate system: Preferably, the measurement output information of the accelerometer in the arrow body coordinate system is calculated as ； wherein: , , , , and are the coefficients of the linear terms of the accelerometers, , , are the outputs of the axis, axis and axis accelerometers, respectively; , , are the coefficients of the zeroth order terms of the axis, axis and axis accelerometers, respectively; is the computation period of the flight control system software; , and are the accelerometer output information in the axis, axis and axis directions after error compensation; the zeroth order error refers to the output voltage of the accelerometer which is not zero when the input is zero due to manufacturing error; the linear term error refers to the small angle between the sensitive axis of the accelerometer and the installation axis during the manufacturing process, which is usually obtained by single machine calibration. (3) calculating the deviation between the measurement output of the accelerometer and the instruction value, and applying the mean filter algorithm to filter the deviation to provide more accurate input for control compensation; (4) calculating the correction amount of the acceleration instruction according to the deviation amount between the expected thrust and the actual thrust; (5) obtaining the corresponding thrust instruction direction according to the thrust vector direction defined in step (1).

2. The launch vehicle reusability control command compensation method based on accelerometer output feedback according to claim 1, characterized in that Step (3) adopts the method of taking the average value of the accelerometer measurement output in step (2) to perform the following filtering calculation: ； wherein: N is the number of data averaged; subscript respectively represent , or axis; is the accelerometer measurement output value after error compensation in step (2); is the expected acceleration value, which can be derived by differentiating the flight trajectory planned by the user.

3. The launch vehicle reusability control command compensation method based on accelerometer output feedback according to claim 1, characterized in that Step (4) Corrects the acceleration command according to the deviation amount of the desired thrust from the actual thrust is calculated as ； where: is the current mass of the aircraft.

4. The launch vehicle reusability control command compensation method based on accelerometer output feedback according to claim 1, characterized in that Step (5) is to derive a corrected thrust command from step (4) To ； In the formula: is the coefficient of compensation quantity, which can be selected according to the characteristics of the engine, and is usually taken as 0.5; is the gravitational acceleration of the aircraft at the current time, and the calculation formula is ； where: , , , , , , , ; G is the Earth gravitational constant, with a value of 3.986005 x 1014m 3 / s 2 ; J2 is the Earth J2 term gravitational coefficient, with a value of 0.00108263; R is the Earth equatorial radius, with a value of 6378140 m; is the launch azimuth; is the launch point geodetic latitude; , and are the 3-axis components of the position of the Earth center to the launch point in the launch coordinate system; , and are the 3-axis components of the vehicle in the launch coordinate system.

5. By the modified thrust command The expression for the desired thrust magnitude is ； According to the coordinate system definition in step (1), the expected attitude angle of the thrust vector direction is obtained as ； wherein and the value ranges of the variables a, b, c, d, e, f, g, h, i, j, k ； The desired thrust is sent to the power system to produce a corresponding thrust, the thrust direction , is sent to the attitude control system to produce a corresponding thrust vector.

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