Aircraft attitude control method based on single-frame control moment gyro servo

Abstract

The invention provides an aircraft attitude control method based on single-frame control moment gyroscope servo, which comprises the following steps: designing a single-frame control moment gyroscope (SGCMG) control rate, and obtaining a frame angular rate according to a moment instruction; according to the moment instruction and the frame angular rate, the moment of the single-frame control moment gyroscope acting on the aircraft is obtained; according to the moment instruction, obtaining a frame angular rate and a moment acting on the aircraft by a single-frame control moment gyroscope, and constructing an aircraft attitude dynamics model and an attitude kinematics equation model; according to the aircraft attitude dynamics model and the attitude kinematics equation model, control law parameters are designed, and an attitude control law equation is obtained. In order to solve the problems that a high-speed aircraft executing mechanism needs to bear long-time high temperature resistance and the control moment of a yaw / rolling channel is insufficient, a single-frame control moment gyroscope SGCMG is adopted as an internal executing mechanism, an aircraft attitude control system is designed, and attitude stability and attitude maneuvering of a high-maneuvering aircraft are achieved.

Description

Technical Field

[0001] The invention belongs to the technical field of aircraft navigation, guidance and control, and in particular relates to an aircraft attitude control method based on a single-frame control moment gyro servo. Background Art

[0002] High-speed aircraft have the characteristics of high speed and short reaction time. They generally use aerodynamic surfaces, variable center of mass, RCS, etc. as attitude control actuators. However, aerodynamic surfaces have problems such as high-temperature ablation and low control efficiency. Although variable center of mass control does not have the problem of ablation, it is highly dependent on aerodynamic forces, and the movement of the mass slider alone cannot make the aircraft produce large attitude maneuvers, and the design of the attitude controller is relatively complex. The Rejection Control System (RCS) can make up for the lack of low control efficiency of aerodynamic surfaces, and can provide constant control torque, and also has the advantage of fast response speed. However, RCS needs to consume fuel, which will occupy a large space inside the aircraft, and RCS generally uses a switch-type nozzle, which is difficult to achieve precise attitude control. In addition, the jet of RCS will interfere with the surface flow field of the aircraft.

[0003] In view of the problems that the actuators of high-speed aircraft need to withstand long-term high temperature, the yaw / roll channel control torque is insufficient, the large internal space of the aircraft is occupied, or it is difficult to achieve precise attitude control, it is necessary to provide a new aircraft attitude control method. The single gimbal control moment gyroscope (SGCMG) can directly output torque to control the attitude of the aircraft. It does not need to contact with the external high-temperature fluid like the air rudder, so it greatly reduces the demand for high-temperature resistant design; when the aircraft enters the outer space of the atmosphere, the aircraft attitude can continue to be controlled without the problem of air rudder failure. In addition, the SGCMG can be arranged at any position of the aircraft, which provides convenience for the weight balancing of the aircraft. However, at present, high-speed weapon aircraft have not yet applied the single-frame control moment gyro servo as a control stand-alone product, so it is urgent to carry out research on the aircraft attitude control method based on the single-frame control moment gyro servo. Summary of the invention

[0004] In order to overcome the deficiencies in the prior art, the inventors have conducted intensive research and provided an aircraft attitude control method based on a single-frame control moment gyro servo. The single-frame control moment gyro SGCMG is used as an internal actuator to design an aircraft attitude control system to achieve attitude stabilization and attitude maneuverability of a highly maneuverable aircraft.

[0005] The technical solution provided by the present invention is as follows:

[0006] In a first aspect, an aircraft attitude control method based on a single-frame control moment gyro servo comprises:

[0007] Design the control rate of the single frame control moment gyro SGCMG, and obtain the frame angular rate according to the torque command;

[0008] According to the torque command and the frame angular rate, the torque of the single frame control moment gyro acting on the aircraft is obtained;

[0009] According to the torque command, the frame angular rate and the torque of the single-frame control moment gyro acting on the aircraft are obtained, and the aircraft attitude dynamics model and attitude kinematics equation model are constructed;

[0010] According to the aircraft attitude dynamics model and attitude kinematics equation model, the control law parameters are designed and the attitude control law equation is obtained.

[0011] In a second aspect, an aircraft attitude control device based on a single-frame control moment gyro servo comprises:

[0012] one or more processors;

[0013] a storage device for storing one or more programs,

[0014] When the one or more programs are executed by the one or more processors, the one or more processors implement the aircraft attitude control method based on single-frame control moment gyro servo as described in the first aspect.

[0015] In a third aspect, a readable storage medium stores a computer program, which, when executed by a processor, implements the aircraft attitude control method based on single-frame control moment gyro servo as described in the first aspect.

[0016] In a fourth aspect, a computer program product is provided, comprising: a computer program (also referred to as code, or instruction), which, when executed, executes the aircraft attitude control method based on single-frame control moment gyro servo described in the first aspect.

[0017] The invention provides an aircraft attitude control method based on a single-frame control moment gyro servo, which has the following beneficial effects:

[0018] (1) The present invention provides an aircraft attitude control method based on a single-frame control moment gyro servo, and for the first time proposes a scheme using a control moment gyro as a three-axis attitude control actuator. The scheme has the advantages of large control torque, no ablation problem, and fast response speed. At the same time, it has high working efficiency and good dynamic characteristics, and is more suitable for attitude stabilization and control of highly maneuverable aircraft.

[0019] (2) The present invention provides an aircraft attitude control method based on a single-frame control moment gyro servo, establishes an aircraft dynamics model that takes into account the characteristics of the control moment gyro, and lays a foundation for the application of attitude control technology in high-speed aircraft based on control moment gyro;

[0020] (3) The present invention provides an aircraft attitude control method based on a single-frame control torque gyro servo, and proposes an attitude control scheme based on a control torque gyro, which solves the shortcoming of insufficient control torque of the yaw / roll channel of the air rudder. Mathematical simulation results show that in the presence of initial condition errors and external interference torques, the aircraft can track the instructions well with a small error. The theoretical method can be applied to practical engineering. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] Figure 1 The flowchart of the aircraft attitude control method based on the single frame control moment gyro servo;

[0022] Figure 2 is the yaw attitude angle tracking curve;

[0023] Figure 3 is the yaw attitude angle tracking error curve;

[0024] Figure 4 is the yaw rate curve;

[0025] Figure 5 This is the torque curve of the torque gyro control. DETAILED DESCRIPTION

[0026] The following detailed description of the present invention will make the features and advantages of the present invention more clear and explicit.

[0027] The word “exemplary” is used exclusively herein to mean “serving as an example, example, or illustration.” Any embodiment described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other embodiments.

[0028] The present invention provides an aircraft attitude control method based on a single-frame control moment gyro servo, comprising the following steps:

[0029] Step 1: Design the single-frame control moment gyro SGCMG control rate and obtain the frame angular rate according to the torque command

[0030] The control law of SGCMGs refers to the given torque command M cmd Solve for the angular velocity of each frame The present invention uses a singular avoidance escape control law, which has a smaller torque tracking error, a lower frame angular velocity and a faster escape velocity. The formula is as follows:

[0031]

[0032] where F ° ∈R n×m is the filtered singular escaped pseudo-inverse matrix, is the singular avoidance angular velocity, M cmd =[CMX cmd , CMY cmd CMZ cmd ] is the torque command, I n is the unit matrix, μ is the filter factor, and F is the set of unit axial quantities of the frame coordinate system.

[0033] Step 2: According to the torque command and frame angular velocity The torque of the single-frame control moment gyro acting on the aircraft is obtained.

[0034] The dynamic equation of the single-frame control moment gyro is:

[0035]

[0036] Among them, the inertia matrix Ф is:

[0037]

[0038] In the kinetic equation, J s is the total moment of inertia of the flywheel and frame, J Tg is the moment of inertia of the frame in the direction of frame rotation, J Wh is the moment of inertia of the flywheel in the direction of flywheel rotation, ω∈R 3 is the angular velocity of the aircraft in the inertial coordinate system, γ∈R n is the frame angle, Ω∈R n is the flywheel angular velocity. M g ∈R n Input torque command for frame motor, M h ∈R n is the flywheel motor input torque, M ext ∈R 3 M is the disturbance torque of the aircraft in the inertial coordinate system. ω 、M γ 、M Ω is the torque generated by SGCMGs on the aircraft, moment gyro frame and flywheel due to the aircraft angular velocity, frame angular velocity and flywheel angular velocity. The specific calculation is as follows:

[0039]

[0040] M γ =diag(F Tω)((J Th -J Tf )H T ω+J Wh Ω) (5)

[0041]

[0042] Among them, J Tf is the inertia of the flywheel and frame combination in the torque direction, J Tg is the inertia of the flywheel and frame combination in the direction of frame rotation, J Th is the inertia of the flywheel and frame combination in the direction of flywheel rotation, and F, G, and H are the sets of unit axial quantities of the frame coordinate system:

[0043] F=[f1…f n ] (7)

[0044] G=[g1…g n ] (8)

[0045] H=[h1…h n ] (9)

[0046] f i is the output torque unit vector of the i-th control torque gyro, g i is the unit vector of the rotation direction of the i-th control torque gyro frame, h i is the unit vector of the rotation direction of the i-th control torque tourbillon;

[0047] The vehicle achieves attitude control by exchanging angular momentum with the SGCMGs. When the SGCMGs are subjected to the torque M int When the aircraft is subjected to -M int The reaction torque of the single-frame control moment gyro can be obtained from the dynamic equation:

[0048]

[0049] Assume that the inertia of the moment gyro frame is J T Much smaller than the vehicle's inertia tensor, J T ≈J B , J B is the inertia tensor of the aircraft; the flywheel rotation speed is constant at Ω c .,Right now Frame angular acceleration is also very small, then the control torque generated by SGCMGs is:

[0050]

[0051] Where τ is the control torque generated by SGCMGs due to the frame angular velocity, Md is the disturbance torque generated by SGCMGs due to the angular velocity of the aircraft. Then the real SGCMGs control command should be:

[0052] τ ref =M int -M d (12)

[0053] From formulas (2) and (5), the frame torque command can be obtained as:

[0054] M cmd =M g ≈-FωJ Wh Ω c (13)

[0055] Because the angular velocity of the aircraft is much smaller than the angular velocity of the frame,

[0056] Then the torque magnification of SGCMGs is:

[0057]

[0058] It can be seen that the torque amplification factor of SGCMGs is proportional to the ratio of the frame angular velocity and the vehicle angular velocity, so SGCMGs can produce a larger torque amplification effect.

[0059] Step 3, according to the torque command, obtain the frame angular rate and the torque of the single-frame control moment gyro acting on the aircraft, and construct the aircraft attitude dynamics model and attitude kinematics equation model.

[0060] The aircraft attitude dynamics equation is constructed as follows:

[0061]

[0062] Where: J x1 , J y1 , J z1 are the moments of inertia of the aircraft in the x, y, and z axes in the body coordinate system; M x1 、M y1 、M z1 are the aerodynamic moments in the x, y, and z directions respectively; M x_int 、M y_int 、M z_int are the torques in the x, y, and z directions generated by the control torque gyro; ω x1 ,ω y1 ,ω z1 They are respectively the roll angular rate, yaw angular rate, and pitch angular rate solved by the aircraft.

[0063] The aircraft attitude kinematic equation is as follows

[0064]

[0065] Where: γ, ψ, They are the roll angle, yaw angle and pitch angle in the inertial system respectively.

[0066] Step 4, design attitude control law: Based on the aircraft attitude dynamics model and attitude kinematics equation model obtained in step 3, design the control law parameters and obtain the attitude control law equation.

[0067] The attitude control law equation is as follows:

[0068]

[0069] Tracking error is defined as

[0070] Where: γ cmd , cmd , are the desired roll angle, yaw angle, and pitch angle respectively; ω x1 ,ω y1 ,ω z1 They are respectively the roll angular rate, yaw angular rate, and pitch angular rate solved on the missile; are the control gains of pitch, yaw and roll channels respectively; It is the dynamic gain of the pitch channel; is the dynamic gain of the yaw channel; is the dynamic gain of the roll channel; is the gain of the integral term of the pitch / yaw / roll channel; These are the roll / yaw / pitch channel attitude control networks that need to be designed; CMX cmd , CMY cmd CMZ cmd They are the torque commands for the roll channel, yaw channel, and pitch channel respectively.

[0071] Figure 2-Figure 5 This is a simulation verification result diagram of the present invention taking the yaw channel as an example. It can be seen that the attitude angle errors are all within 0.18°, and the angular rate converges within 1° / s. The control method proposed in the present invention has good control accuracy.

[0072] Step 5, design the attitude controller correction network and parameter selection criteria;

[0073] Three-channel attitude control network The form is as follows:

[0074]

[0075] Wherein ω1, ω2, ω3, ξ1, ξ3 are the parameters that need to be designed, ω1, ω2, ω3 are selected near the aircraft cutoff frequency, and ξ1, ξ3 are selected between 0 and 1.

[0076] The present invention also provides an aircraft attitude control device based on a single-frame control moment gyro servo, comprising:

[0077] one or more processors;

[0078] a storage device for storing one or more programs,

[0079] When the one or more programs are executed by the one or more processors, the one or more processors implement the aircraft attitude control method based on single-frame control moment gyro servo as described in the first aspect.

[0080] The present invention also provides a readable storage medium having a computer program stored thereon, which, when executed by a processor, implements the aircraft attitude control method based on single-frame control moment gyro servo as described in the first aspect.

[0081] The readable storage medium includes, but is not limited to, various media that can store program codes, such as a USB flash drive, a mobile hard disk, a read-only memory (ROM), a random access memory (RAM), a magnetic disk or an optical disk.

[0082] The present invention also provides a computer program product, which includes: a computer program (also referred to as code, or instruction), which, when executed, executes the aircraft attitude control method based on single-frame control moment gyro servo described in the first aspect.

[0083] In the above embodiments, it can be implemented in whole or in part by software, hardware, firmware or any combination thereof. When implemented using software, it can be implemented in whole or in part in the form of a computer program product. The computer program product includes one or more computer instructions. When the computer instructions are loaded and executed on a computer, the process or function described in the embodiment of the present application is generated in whole or in part. The computer can be a general-purpose computer, a special-purpose computer, a computer network, or other programmable device. The computer instructions can be stored in a computer-readable storage medium, or transmitted from one computer-readable storage medium to another computer-readable storage medium, for example, the computer instructions can be transmitted from a website site, computer, server or data center by wired (e.g., coaxial cable, optical fiber, digital subscriber line (digital subscriber line, DSL)) or wireless (e.g., infrared, microwave, etc.) mode to another website site, computer, server or data center.

[0084] Those of ordinary skill in the art will appreciate that the units and algorithm steps of each example described in conjunction with the embodiments disclosed herein can be implemented in electronic hardware, or a combination of computer software and electronic hardware. Whether these functions are performed in hardware or software depends on the specific application and design constraints of the technical solution. Professional and technical personnel can use different methods to implement the described functions for each specific application, but such implementation should not be considered to be beyond the scope of this application.

[0085] Those skilled in the art can clearly understand that, for the convenience and brevity of description, the specific working processes of the above-described equipment, devices and modules can refer to the corresponding processes in the aforementioned method embodiments and will not be repeated here.

[0086] The present invention has been described in detail above in conjunction with specific implementations and exemplary examples, but these descriptions cannot be understood as limiting the present invention. Those skilled in the art understand that, without departing from the spirit and scope of the present invention, a variety of equivalent substitutions, modifications or improvements may be made to the technical solution of the present invention and its implementation methods, all of which fall within the scope of the present invention. The scope of protection of the present invention shall be subject to the attached claims.

[0087] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

Claims

1. A method for aircraft attitude control based on single frame control moment gyro servo, characterized in that: include: Design the control rate of the single frame control moment gyro SGCMG, and obtain the frame angular rate according to the torque command; According to the torque command and the frame angular rate, the torque of the single frame control moment gyro acting on the aircraft is obtained; According to the torque command, the frame angular rate and the torque of the single-frame control moment gyro acting on the aircraft are obtained, and the aircraft attitude dynamics model and attitude kinematics equation model are constructed; According to the aircraft attitude dynamics model and attitude kinematics equation model, the control law parameters are designed and the attitude control law equation is obtained.

2. The aircraft attitude control method based on single frame control moment gyro servo according to claim 1, characterized in that: In the step of designing the single-frame control moment gyro SGCMG control rate and obtaining the frame angular rate according to the torque command, the frame angular rate is determined by the following formula: in, is the frame angular velocity, F v ∈R n×m is the filtered singular escaped pseudo-inverse matrix, is the singular avoidance angular velocity, M cmd =[CMX cmd , CMY cmd ,CMZ cmd ] is the torque command, I n is the unit matrix, μ is the filter factor, and F is the set of unit axial quantities of the frame coordinate system.

3. The aircraft attitude control method based on single frame control moment gyro servo according to claim 2, characterized in that: In the step of obtaining the torque of the single-frame control moment gyro acting on the aircraft according to the torque command and the frame angular rate, Among them, M int The torque acting on the aircraft by the single-frame control moment gyro; is the frame angular rate, ω∈R 3 is the angular velocity of the aircraft in the inertial coordinate system.

4. The aircraft attitude control method based on single frame control moment gyro servo according to claim 3, characterized in that: In the step of obtaining the frame angular rate and the torque of the single-frame control moment gyro acting on the aircraft according to the torque command, and constructing the aircraft attitude dynamics model and the attitude kinematics equation model, the aircraft attitude dynamics equation is as follows: Where: J x1 , J y1 , J z1 are the moments of inertia of the aircraft in the x, y, and z axes in the body coordinate system; M x1 、M y1 、M z1 are the aerodynamic moments in the x, y, and z directions respectively; M x_int 、M y_int 、M z_int are the torques in the x, y, and z directions generated by the control torque gyro; ω x1 ,ω y1 ,ω z1 They are respectively the roll angular rate, yaw angular rate, and pitch angular rate solved by the aircraft.

5. The aircraft attitude control method based on single frame control moment gyro servo according to claim 4, characterized in that: In the step of obtaining the frame angular rate and the torque of the single-frame control moment gyro acting on the aircraft according to the torque command, and constructing the aircraft attitude dynamics model and attitude kinematics equation model, the aircraft attitude kinematics equation is as follows Where: γ, ψ, They are the roll angle, yaw angle and pitch angle in the inertial system respectively.

6. The aircraft attitude control method based on single frame control moment gyro servo according to claim 1, characterized in that: In the step of designing control law parameters and obtaining attitude control law equations according to the aircraft attitude dynamics model and attitude kinematics equation model, the attitude control law equation is as follows: Tracking error is defined as Where: γ cmd , cmd , are the desired roll angle, yaw angle, and pitch angle respectively; ω x1 ,ω y1 ,ω z1 They are respectively the roll angular rate, yaw angular rate, and pitch angular rate solved on the missile; are the control gains of pitch, yaw and roll channels respectively; It is the dynamic gain of the pitch channel; is the dynamic gain of the yaw channel; is the dynamic gain of the roll channel; is the gain of the integral term of the pitch / yaw / roll channel; These are the roll / yaw / pitch channel attitude control networks that need to be designed; CMX cmd , CMY cmd CMZ cmd They are the torque commands for the roll channel, yaw channel, and pitch channel respectively.

7. The aircraft attitude control method based on single frame control moment gyro servo according to claim 6, characterized in that: The three-channel attitude control network The form is as follows: Wherein ω1, ω2, ω3, ξ1, ξ3 are the parameters that need to be designed, ω1, ω2, ω3 are selected near the aircraft cutoff frequency, and ξ1, ξ3 are selected between 0 and 1.

8. An aircraft attitude control device based on a single-frame control moment gyro servo, characterized in that: include: one or more processors; a storage device for storing one or more programs, When the one or more programs are executed by the one or more processors, the one or more processors implement the aircraft attitude control method based on single-frame control moment gyro servo as described in one of claims 1 to 7.

9. A readable storage medium, characterized in that: A computer program is stored thereon, and when the program is executed by a processor, the aircraft attitude control method based on a single-frame control moment gyro servo as claimed in any one of claims 1 to 7 is implemented.

10. A computer program product, characterized in that The computer program product comprises: a computer program, which, when being executed, executes the aircraft attitude control method based on single-frame control moment gyro servo according to any one of claims 1 to 7.

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