New method for active control of global vibration of helicopter body

Abstract

The invention discloses a new method for active control of global vibration of a helicopter body, and provides a method for active control of global vibration of a helicopter body by installing a piezoelectric laminated actuator on a main reducing support rod to control the vibration load of a rotor wing in order to solve the problems that the global vibration control effect of the helicopter body is poor and the transmission of the vibration load of the rotor wing to a helicopter body structure cannot be inhibited by a traditional structural vibration active control method. The invention relates to a novel active control method for global vibration of a helicopter body with a stay bar vibration load as a control target. The purpose of reducing the global vibration level of the helicopter body is achieved. The method can suppress the rotor vibration load transmitted to the helicopter body structure, has the advantage of efficiently suppressing the global vibration level of the helicopter body, and has the control convergence speed superior to that of a traditional method.

Description

Technical Field

[0001] The invention belongs to the technical field of helicopter vibration control. Background Art

[0002] During helicopter flight, the rotor vibration load is transmitted to the fuselage structure through the main struts, causing severe fuselage vibration response. Traditional helicopter vibration passive control methods use technologies such as dynamic vibration absorption and vibration isolation to achieve vibration control without external energy input. This type of method has high reliability and simple structure, but has disadvantages such as large installation mass, limited control effect, narrow control frequency band and poor adaptability, and can no longer meet the increasing requirements for reducing helicopter vibration levels.

[0003] The active vibration control method of helicopters can effectively overcome the defects of passive control technology and has the potential and broad prospects for achieving efficient control of helicopter vibration levels. According to different control objectives, the active vibration control methods of helicopters can be roughly divided into two categories: active vibration control methods based on rotor systems and active vibration control methods based on fuselage structures. The active vibration control method based on the rotor system adjusts the unsteady aerodynamic load of the blades by applying active control force to the rotor system, thereby reducing the vibration load superimposed on the hub. It mainly includes high-order harmonic control, independent blade control, trailing edge winglet control and active torsion rotor. This method can effectively suppress the rotor vibration load, but because the actuation system is arranged in the rotor system, the control power requirement is high and it will affect the airworthiness. In addition, how to realize the transfer of control energy to the rotor system is a difficult problem faced by this method. The active vibration control method based on the fuselage structure arranges the actuator and sensor in the fuselage structure to drive the fuselage structure to generate an actuation response, which offsets the excitation response caused by the rotor vibration load at the sensor, thereby suppressing the vibration at the target measurement point, which is called structural response active control. Compared with the active vibration control method based on the rotor system, this method has the advantages of better control effect, lower control power requirement, easier implementation, and no impact on airworthiness. However, this method takes the response of the target measuring point as the control target and is a local vibration control. Limited by the number of vibration response sensors, the active structural response control cannot effectively suppress the global vibration response of the helicopter body, and may even amplify the vibration level outside the target measuring point. Summary of the invention

[0004] Purpose of the invention: In order to solve the problems existing in the above-mentioned background technology, the present invention provides a new method for active control of the global vibration of a helicopter body.

[0005] Technical solution: The present invention provides a new method for active control of the global vibration of a helicopter body, which specifically includes the following steps: S1: Two piezoelectric stack actuators are arranged in parallel on a main strut rod to form a group of active struts, and both ends of each piezoelectric stack actuator are installed on the main strut rod through fasteners and elastic hinges; a force sensor is arranged at one end of the main strut rod connected to the machine body to measure the vibration load of the strut rod; S2: Before the control starts, the steady-state driving voltage at the blade passing frequency is input to each active strut, and the vibration load response of all struts is collected; the frequency response function from each active strut to the strut load is calculated, and the system frequency response function matrix is ​​generated; S3: collecting the vibration load response of each main strut through a force sensor as a control error response, and extracting the harmonic coefficient of the control error response; S4: using the system frequency response function matrix generated in S2 and the control error response harmonic coefficient of the previous moment extracted in S3, update the harmonic coefficient of the control input at the next moment; S5: Using the harmonic coefficients of the control input generated in S4, the time domain signal of the control input at the next moment is calculated, and after analog / digital conversion and power amplifier processing, it is input to the piezoelectric stack actuator of each active strut to generate actuating force, and then transferred to S3.

[0006] Furthermore, the system frequency response function matrix in S2 is specifically:

[0007] Where T is the system frequency response function matrix, M and K are the number of input and output signals of the active vibration control system, is the frequency response from the mth input to the kth output at the blade passing frequency ω, Re represents the real part of the complex number, and Im represents the imaginary part of the complex number.

[0008] Furthermore, the specific method for extracting the control error response harmonic coefficient in S3 is:

[0009] in, is the control error response harmonic coefficient at the nth sampling moment, is the control error response vector at the nth sampling moment, is the harmonic identification step size, is the harmonic basis function vector, is the sampling interval.

[0010] Furthermore, the S4 is specifically: S41: Establish a quadratic control objective function regarding control input and control error response:

[0011] in, is the harmonic coefficient of the control input, and are the weight coefficient matrices of the control error response and control input respectively; S42: Set the control objective function to 0, and obtain the update formula of the control input harmonic coefficient; update the control input harmonic coefficient at the next moment according to the system frequency response function matrix, the control error response harmonic coefficient at the previous moment, and the control input harmonic coefficient.

[0012] Beneficial effects: The present invention realizes the global control of helicopter body vibration under the excitation of rotor vibration load, effectively suppresses the rotor vibration load transmitted to the body structure by the main strut, avoids the defect of insufficient response control of the structure response active control to the position outside the target measuring point, and solves the problem that the existing control method cannot control the global vibration of the helicopter body. The example shows that the vibration of the global 20 measuring points of the body structure is well controlled by adopting the new method of active control of the global vibration of the helicopter body of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0013] Figure 1 is a system block diagram of the present invention; Figure 2 It is a schematic diagram of the installation of the active strut; Figure 3 It is the response amplitude of 20 measuring points of the helicopter body under the same rotor vibration load excitation, respectively using this control method, the traditional structural response active control method and no control, among which measuring points 1-6 are the target measuring points of the structural response active control; Figure 4 is the square root of the sum of the squares of the helicopter body target measurement point and 20 measurement points when the control method and the traditional structural response active control method are used under the same rotor vibration load excitation; Figure 4 (a) is the square root of the sum of the squares of the helicopter body target point and 20 measurement points when the traditional structural response active control method is used. Figure 4 (b) is the square root of the sum of the squares of the responses of the helicopter body target measurement point and the 20 measurement points when this control method is used. DETAILED DESCRIPTION

[0014] The accompanying drawings, which constitute a part of the present invention, are used to provide a further understanding of the present invention. The exemplary embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute improper limitations on the present invention.

[0015] like Figure 1 The embodiment shown provides a new method for active global vibration control of a helicopter body, including the following steps: S1: Two piezoelectric stack actuators are arranged in parallel on a main strut rod to form a group of active struts, and both ends of each piezoelectric stack actuator are installed on the main strut rod through fasteners and elastic hinges; a force sensor is arranged at one end of the main strut rod connected to the machine body to measure the vibration load of the strut rod; S2: Before the control starts, the steady-state driving voltage at the blade passing frequency is input to each active strut, and the vibration load response of all struts is collected; the frequency response function from each active strut to the strut load is calculated, and the system frequency response function matrix is ​​generated; S3: collecting the vibration load response of each main strut through a force sensor as a control error response, and extracting the harmonic coefficient of the control error response; S4: using the system frequency response function matrix generated in S2 and the control error response harmonic coefficient of the previous moment extracted in S3, update the harmonic coefficient of the control input at the next moment; S5: Using the harmonic coefficients of the control input generated in S4, the time domain signal of the control input at the next moment is calculated, and after analog / digital conversion and power amplifier processing, it is input to the piezoelectric stack actuator of each active strut to generate actuating force, and then transferred to S3.

[0016] In this embodiment, S1 specifically comprises: arranging two piezoelectric stack actuators in parallel on a main support rod to form a group of active support rods, such as Figure 2 As shown, one end of the piezoelectric stack actuator is connected to the rigid connection element through a bolt, and the other end is connected to the rigid connection element through an elastic hinge and a bolt. The two rigid connection elements are fastened to both sides of the support rod by bolts, wherein the elastic hinge is used to prevent the piezoelectric stack actuator from being damaged by excessive bending load; a force sensor is arranged at one end of the main support rod connecting the machine body to measure the vibration load of the support rod; In this embodiment, S2 is specifically: before the control starts, a steady-state driving voltage at the blade passing frequency is input to each active strut. , collect the vibration load response of all struts ; Use Fourier transform to calculate the driving voltage in the complex frequency domain and strut vibration load response , where j is an imaginary unit, m represents the mth active strut, and k represents the kth strut vibration load. The frequency response function of each active strut to strut load is calculated: , generate the system frequency response function matrix: (1) Where Re represents the real part of the complex number, and Im represents the imaginary part of the complex number.

[0017] In this embodiment, S3 specifically includes: collecting the vibration load response of each main deflector rod through the force sensor as the control error response , using the minimum mean square error method, the control error response harmonic coefficient at the current sampling moment is extracted based on the control error response and the control error response harmonic coefficient at the previous sampling moment: (2) in, is the control error response harmonic coefficient at the nth sampling moment, is the control error response vector at the nth sampling moment, is the harmonic identification step size, is the harmonic basis function vector, is the sampling interval.

[0018] In this example, S4 can be divided into two steps, where: Step A: Establish a quadratic control objective function regarding control input and control error response: (3) in, is the harmonic coefficient of the control input, and are the weight coefficient matrices of the control error response and control input, respectively, , , represents the weight coefficient of the kth control error response, represents the weight coefficient input to the mth control input; Step B: Set the control objective function to 0 and obtain the update formula of the control input harmonic coefficient: (4) in, , It is the harmonic correction factor, which is used to limit the rate of change of the control input to prevent the control input from being too large.

[0019] The control input harmonic coefficient at the next moment is updated according to the system frequency response function matrix, the control error response harmonic coefficient at the previous moment and the control input harmonic coefficient.

[0020] In this embodiment, S5 is specifically: using the harmonic coefficient of the control input generated in S4 , calculate the time domain signal of the control input at the next moment: (5) The time domain signal of the control input is processed by analog / digital conversion and power amplifier, input to the piezoelectric stack actuator of each active strut to generate actuating force, and transferred to S3.

[0021] Figure 3The response amplitudes of 20 measuring points of the helicopter body under the same rotor vibration load excitation are given when the control method, the traditional structural response active control method and no control are used respectively. The structural response active control effectively suppresses the vibration response of the target measuring point, but the control effect on the responses of the remaining measuring points is general, and even amplifies the vibration response of the measuring point 20. Although the control effect of the vibration response of the target measuring point by the method of the present invention is slightly inferior to that of the structural response active control, the responses of all measuring points can be efficiently suppressed, indicating that the method of the present invention has the advantage of effectively controlling the global vibration level of the helicopter body.

[0022] Figure 4 (a) and (b) in the figure respectively give the square root of the sum of squares of the helicopter body target measurement point response and the square root of the sum of squares of all measurement point responses when the control method and the traditional structural response active control method are used. The control method of the present invention has a slightly lower suppression rate for the target measurement point vibration level than the structural response active control, but can greatly improve the control effect on the global vibration level of the body, indicating that the method of the present invention can effectively suppress the target measurement point and the global vibration level of the body at the same time, and has a faster control convergence speed than the structural response active control.

[0023] The embodiments of the present invention are described in detail above with reference to the accompanying drawings, but the present invention is not limited to the above embodiments, and various changes can be made within the knowledge scope of ordinary technicians in this field without departing from the purpose of the present invention.

Claims

1. A new method for active control of global vibration of helicopter body, characterized in that: The specific steps include: S1: Two piezoelectric stack actuators are arranged in parallel on a main strut rod to form a group of active struts, and both ends of each piezoelectric stack actuator are installed on the main strut rod through fasteners and elastic hinges; a force sensor is arranged at one end of the main strut rod connected to the machine body to measure the vibration load of the strut rod; S2: Before the control starts, the steady-state driving voltage at the blade passing frequency is input to each active strut, and the vibration load response of all struts is collected; the frequency response function from each active strut to the strut load is calculated, and the system frequency response function matrix is ​​generated; S3: collecting the vibration load response of each main strut through a force sensor as a control error response, and extracting the harmonic coefficient of the control error response; S4: using the system frequency response function matrix generated in S2 and the control error response harmonic coefficient of the previous moment extracted in S3, update the harmonic coefficient of the control input at the next moment; S5: Using the harmonic coefficients of the control input generated in S4, the time domain signal of the control input at the next moment is calculated, and after analog / digital conversion and power amplifier processing, it is input to the piezoelectric stack actuator of each active strut to generate actuating force, and then transferred to S3.

2. The helicopter body global vibration control method with active control of the main strut vibration load according to claim 1 is characterized in that: The system frequency response function matrix in S2 is specifically: Where T is the system frequency response function matrix, M and K are the number of input and output signals of the active vibration control system, is the frequency response from the mth input to the kth output at the blade passing frequency ω.

3. The helicopter body global vibration control method with active control of main strut vibration load according to claim 1 is characterized in that: The specific method of extracting the control error response harmonic coefficient in S3 is: in, is the control error response harmonic coefficient at the nth sampling moment, is the control error response vector at the nth sampling moment, is the harmonic identification step size, is the harmonic basis function vector, is the sampling interval.

4. The new method for active control of helicopter body global vibration according to claim 1 is characterized in that: The S4 is specifically: S41: Establish a quadratic control objective function regarding control input and control error response: in, is the harmonic coefficient of the control input, and are the weight coefficient matrices of the control error response and control input respectively; S42: Set the control objective function to 0, and obtain the update formula of the control input harmonic coefficient; update the control input harmonic coefficient at the next moment according to the system frequency response function matrix, the control error response harmonic coefficient at the previous moment, and the control input harmonic coefficient.

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