A tiltrotor wing vibration active control system and method

Abstract

The application discloses a tilt-rotor wing vibration active control system and method, which comprises: an actuator assembly composed of piezoelectric laminated actuators symmetrically arranged in the wing; an acceleration sensor comprising a wing acceleration sensor arranged at the outer end of the wing and a fuselage acceleration sensor arranged in the cockpit; and an adaptive control unit for collecting sensor signals and outputting control instructions. Double-loop control and flight state tracking strategies are adopted: the first control loop implements dynamic characteristic control on the wing-rotor system to suppress transient response; and the second control loop implements steady-state response control on the aircraft structure to suppress steady-state harmonic vibration of the rotor passing frequency and its integer multiples. According to the rotor speed and the nacelle tilt angle, the flight mode is distinguished, and the double-loop control weight is dynamically adjusted. The application realizes adaptive vibration suppression of the tilt-rotor aircraft within the full flight envelope, takes into account transient and steady-state response control, and improves flight quality and the service life of key equipment.

Description

Technical Field

[0001] This invention belongs to the field of vibration control technology, and in particular relates to an active control system and method for the vibration of a tiltrotor wing. Background Technology

[0002] Tiltrotor configurations can switch between helicopter and fixed-wing flight modes, possessing unique flight capabilities and versatility, making them an important force in the aviation field. High-level vibrations caused by rotor aerodynamic imbalances and rotor-wing interference, resulting in aeroelastic coupling, are significant factors affecting the flight safety and comfort of tiltrotor configurations. Several active vibration control systems have been implemented abroad to reduce the vibration levels of tiltrotor configurations, achieving good results and meeting the vibration environment requirements of both humans and equipment. Therefore, tiltrotor vibration analysis and control is an important topic and task in the field of tiltrotor configuration research.

[0003] Current research on vibration control strategies for tiltrotor configurations mainly includes two methods: active and passive. Passive methods include aerodynamic layout optimization, structural dynamics optimization, and rotor hub bilinear oscillation vibration absorption. Active control methods include high-order harmonic control of the rotor system, active flap control, and mechanical actuator-driven airframe structure vibration control. Overall, vibration control strategies for tiltrotor configurations mainly rely on mature helicopter vibration reduction technologies. Active flap control developed for tiltrotor configurations also has problems such as coupling with the flight control system and low efficiency of high-frequency drive. Summary of the Invention

[0004] The purpose of this invention is to provide an active control system and method for tiltrotor wing vibration. By employing piezoelectric laminated actuators to drive the active wing, a dual-loop control strategy, and a recursive optimal harmonic control algorithm, the performance and control capability of the tiltrotor configuration vibration control system can be effectively improved, thereby enhancing the flight safety and comfort of the tiltrotor configuration.

[0005] The technical solution of the present invention: According to a first aspect of the present invention, an active vibration control system for a tiltrotor wing is proposed, comprising: an actuator assembly, an acceleration sensor, and an adaptive control unit; the actuator assembly includes a plurality of piezoelectric laminated actuators, centrally symmetrically arranged within the tiltrotor wing configuration on both sides of the fuselage; this symmetrical arrangement can simultaneously cancel the symmetrical and anti-symmetrical vibration modes of the left and right wings, avoiding the introduction of additional unbalanced loads; the acceleration sensor includes a wing acceleration sensor respectively disposed at the outer end of the tiltrotor wing configuration and a fuselage acceleration sensor inside the cockpit, respectively used to collect tilt... The vibration source signal of the rotor configuration wing and the vibration response at the evaluation point in the cockpit; this hybrid sensing strategy of feedforward and feedback can both detect disturbances transmitted from the rotor in advance and accurately control residual vibrations at key locations; the adaptive control unit is used to collect and condition the signals of the acceleration sensor, and adjust the control parameters online in real time through the adaptive control algorithm (such as the Fx-LMS algorithm) built into the adaptive control unit, and output control commands to the internal piezoelectric drive power supply, which outputs high-voltage control energy to the actuator assembly, thereby realizing adaptive vibration suppression for time-varying speed and flight state.

[0006] The core of piezoelectric stacked actuators lies in the piezoelectric effect, which is the generation of electric charge (positive piezoelectric effect) in certain materials under mechanical stress or deformation (inverse piezoelectric effect) under an electric field. This effect is commonly found in crystalline materials such as quartz and barium titanate, as well as ceramic materials such as lead zirconate titanate (PZT). Among these, PZT ceramics are the preferred choice for manufacturing piezoelectric stacked actuators due to their high piezoelectric coefficient and stability.

[0007] Piezoelectric stacked actuators are composed of multiple layers of piezoelectric ceramic sheets stacked together, connected by electrodes. When a voltage is applied, each ceramic sheet undergoes a slight deformation due to the inverse piezoelectric effect. The stacked layers together generate a significant displacement or force output, the output characteristics of which mainly depend on material properties, the number of stacked layers, and structural design. This structural design allows the actuator to output active control forces of hundreds of kilograms or more while maintaining a compact size, making it highly suitable for high-frequency, high-energy vibration control of aerospace structures.

[0008] In one possible embodiment, the tiltrotor configuration wing has an internal spars structure, which includes a front spars and a rear spars arranged parallel to each other along the span of the fuselage; multiple stringers are distributed parallel to each other between the front and rear spars; the piezoelectric laminated actuators are arranged in pairs between adjacent stringers at the upper and lower edges of the front spars and the upper and lower edges of the rear spars. This arrangement makes full use of the internal space of the wing box, embedding the actuators between the main load-bearing components, enabling them to directly apply bending and torsional control moments to the main spars, effectively suppressing the first-order bending and torsional vibrations of the wing.

[0009] In one possible embodiment, the actuation output axis of the piezoelectric laminated actuator is parallel to and connected in parallel with the front and rear beams. This co-directional arrangement allows the output forces of multiple actuators to be directly superimposed, forming a powerful tensile or compressive force along the beam axis. It is mainly used to efficiently suppress the vertical bending vibration mode of the wing, and has a simple structure and intuitive control law design.

[0010] In one possible embodiment, the axial direction of the piezoelectric laminated actuator is at an angle to the front or rear beam, the angle being used to generate a torque, to excite or suppress the torsional vibration mode of the wing, and to have a coupling control effect on bending vibration; the actuation output axis of each pair of piezoelectric laminated actuators at the upper and lower edges of the front or rear beams is opposite to the angle between the front and rear beams; the actuation output axis of each pair of piezoelectric laminated actuators at the upper and lower edges of the front and rear beams is opposite to the angle between the front and rear beams.

[0011] This cross-pair, reverse-angled installation method can generate torque, which is specifically used to excite or suppress the torsional vibration mode of the wing, while also having a coupling control effect on bending vibration. Thus, multi-modal decoupling control can be achieved through a single hardware layout, improving the control dimension and efficiency of the system.

[0012] Preferably, under space constraints, the piezoelectric laminated actuator is positioned away from the wing's neutral axis. This maximizes the driving efficiency of the piezoelectric drive unit, efficiently suppressing dynamic load transfer through the wing and responding to wing bending and torsional vibrations. The wing's neutral axis is the axis where stress is zero when the wing bends. The design of positioning the actuator away from the neutral axis essentially utilizes the lever principle, transforming the limited actuator stroke and output into a maximum structural control torque by increasing the physical lever arm, thereby achieving efficient structural response control within the space-constrained wing interior. This is logically consistent with the previously mentioned scheme of placing the actuator on the upper and lower edges of the front / rear spar, representing a deeper mechanical guiding principle for the installation location.

[0013] In one possible embodiment, the fuselage acceleration sensors are arranged at key instruments within the cockpit, at the pilot's position (including foot pedals), and at the cockpit seat; the number of sensors is controlled to be between 8 and 15 to balance the distribution of control points of interest and the complexity of the control system. Key instruments typically refer to airborne equipment that is sensitive to vibration environments or whose failure would seriously affect flight safety and mission completion.

[0014] In one possible embodiment, the wing acceleration sensor arrangement is based on the wing modal characteristics of selective control, and is arranged at the peak of the selected control mode. Optionally, based on the principle of importance of low-order modes for vibration control, three sensors are generally arranged at the wingtip, including two vertical and one chordal. The two vertical sensors are arranged on the fore and aft spars of the wing, respectively, for measuring the wing's vertical bending, chordal bending, and torsional vibration response.

[0015] In one possible embodiment, the adaptive control unit includes a piezoelectric drive power supply, an adaptive controller, and a signal conditioner. The signal conditioner has a digital signal acquisition unit (AD unit) and a filtering unit for acquiring signals from the accelerometer and outputting them to the adaptive controller. The adaptive controller has a digital signal processing unit (DA unit), a control signal output unit, and a control algorithm calculation function for providing the signal samples from the signal conditioner to the control algorithm, which calculates the control output and outputs the control signal via the DA unit. The piezoelectric drive power supply has a signal power amplification function for receiving the control signal from the adaptive control unit, amplifying it, and then driving the actuator assembly to output control force and counteract structural vibration.

[0016] According to a second aspect of the present invention, a method for active control of tiltrotor wing vibration is proposed, employing the aforementioned active control system for tiltrotor wing vibration, comprising the following steps:

[0017] Step 1: Signal Acquisition and Conditioning. The wing acceleration sensor and fuselage acceleration sensor acquire vibration acceleration signals in real time. After being amplified and filtered by the signal conditioner, the signals are input to the adaptive controller. At the same time, the adaptive controller obtains rotor speed signals and nacelle tilt attitude signals from the flight control system.

[0018] Step 2: Flight Status Determination and Control Strategy Decision. The adaptive controller determines the current flight mode based on the acquired rotor speed and nacelle tilt angle, and determines the weight allocation of the dual-loop control strategy; wherein, the dual-loop control strategy includes a first control loop and a second control loop, the first control loop implements dynamic characteristic control for the wing-rotor system, and the second control loop implements steady-state response control for the airframe structure;

[0019] Step 3: Obtain the time-domain vibration signal. The adaptive controller samples the conditioned acceleration signal to obtain the time-domain vibration acceleration signal e(k) at the current sampling time k;

[0020] Step 4: Execute the first control loop (dynamic characteristic control). Based on the flight state determination result, execute the first control loop, using the signal from the wing acceleration sensor as input, and employing a time-domain feedback control algorithm to calculate the wing modal control quantity. The control object of the first control loop mainly targets the low-order modes of the rotor-wing system, including the first-order vertical bending mode, the first-order yaw bending mode, and the first-order torsional mode, used to suppress the vibration response of the non-rotor passing frequency and the transient excitation during tilt transition flight.

[0021] Step 5: Execute the second control loop (steady-state response control). Based on the flight state determination result, execute the second control loop, using the signal from the fuselage acceleration sensor as input, and employing a recursive frequency domain optimal feedforward control algorithm to calculate the steady-state harmonic control quantity;

[0022] Step Six: Control Signal Synthesis and Output. The control quantities of the first control loop obtained in Step Four and the second control loop obtained in Step Five are weighted and fused according to the weights determined in Step Two to obtain the total control input harmonic vector; the time-domain control output is calculated based on the total control input harmonic vector, and the analog control signal is output through the DA unit.

[0023] Step 7: Power Amplification and Actuation. The control signal is amplified by the piezoelectric drive power supply and applied to the piezoelectric multilayer actuator, causing it to output an active force with the opposite phase to the control target, thereby implementing vibration control;

[0024] Step 8: Iterative Loop Control. Measure the vibration response after control in real time using an accelerometer, and repeat steps 1 through 7 to form a closed-loop control until the vibration converges within the target range.

[0025] In one possible embodiment, step two, the flight state determination and control strategy decision, specifically includes:

[0026] a) Determine the flight mode based on the nacelle tilt angle θ: when θ≥75°, it is determined to be helicopter mode; when 75°>θ>15°, it is determined to be tilt transition mode; when θ≤15°, it is determined to be airplane mode.

[0027] b) Determine the weights of the dual-loop control strategy based on the flight mode: In helicopter mode, the first and second control loops have the same weight (W1=W2=0.5); in tilt transition mode, the weight of the first control loop increases (W1=0.7-0.8, W2=0.2-0.3); in airplane mode, the weight of the second control loop increases (W1=0.2-0.3, W2=0.7-0.8).

[0028] c) Determine the target control frequency of the second control loop based on the rotor speed signal Ω, i.e. the rotor passing frequency 1Ω and its integer multiples 2Ω, 3Ω, ..., NΩ, where N is a positive integer;

[0029] d) Based on the nacelle tilt angle signal, switch the weight of the acceleration sensor at different positions, and increase the weight of the wing acceleration sensor in the control target in the tilt transition mode to enhance the suppression of transient response.

[0030] In one possible embodiment, in step four, the time-domain feedback control algorithm used in the first control loop is either a linear quadratic Gaussian (LQG) control algorithm or a modal control algorithm. Using the wing modal coordinates η(t) as the state variable, a state-space equation is constructed. The system state is estimated using a Kalman filter, and the optimal control force is calculated based on the linear quadratic regulator (LQR) to achieve active damping injection of the selected low-order modes, thereby altering the transient response characteristics of the wing-rotor system.

[0031] In one possible embodiment, in step four, the control object of the first control loop is preferentially selected as the low-order mode that has the greatest impact on flight safety and vibration transmission, including the first-order vertical bending mode, the first-order yaw bending mode, and the first-order torsional mode; the control priority of each mode is dynamically adjusted according to the flight mode, for example, in helicopter mode, the focus is on controlling the torsional mode, and in airplane mode, the focus is on controlling the vertical bending mode.

[0032] In one possible embodiment, in step five, the second control loop uses a recursive frequency domain optimal feedforward control algorithm, whose control frequency is updated in real time based on the rotor speed signal; when the rotor speed changes, the algorithm automatically adjusts the frequency components in the reference signal Φ(k) to achieve steady-state harmonic vibration tracking and suppression under variable speed conditions.

[0033] In one possible embodiment, in step six, the fusion of the first control loop and the second control loop adopts a weighted superposition method: the total control quantity u(t) = W1u1(t) + W2u2(t), where u1(t) is the time-domain control quantity output by the first control loop and u2(t) is the time-domain control quantity output by the second control loop; W1 and W2 satisfy W1 + W2 = 1 and are adjusted in real time according to the flight mode.

[0034] In one possible embodiment, step five, which involves calculating the steady-state harmonic control quantity using a recursive frequency domain optimal feedforward control algorithm, includes the following steps:

[0035] The time-domain vibration acceleration signal e(k) is given, where k is the current sampling time.

[0036] Update the control input harmonic coefficient vector using the sampled values ​​of the vibration acceleration signal at the current sampling time. The update formula used for calculation is as follows:

[0037]

[0038]

[0039]

[0040] In the formula, Let k be the acceleration response coefficient vector of the control point at time k. Let K be the acceleration response coefficient vector of the control point at time k-1. , To control the harmonic frequencies, i = 1, 2, ..., R, Forgetting factor, initial gain matrix It can be any positive definite matrix, and can be selected. =I is the identity matrix;

[0041] From the formula Calculate the optimal control input harmonic coefficients, and then update the control input harmonic vector. ,in The transfer matrix from the piezoelectric drive control voltage of the main engine wing to the acceleration response at the vibration observation point is... These are the control point response weighting matrix and the control input weighting matrix, respectively. Let μ be the harmonic coefficient vector of the acceleration response at the control point at time k-1, and μ be the correction step size. and The control input harmonic coefficients for the k-th control correction time step and the previous step;

[0042] Calculate time-domain control output The DA unit outputs control commands;

[0043] The control command is applied to the actuator via a piezoelectric drive power supply to achieve drive force output and implement vibration control.

[0044] According to a third aspect of the present invention, a computer-readable storage medium is provided for storing a computer program, wherein the above-described control method is executed when the computer program is executed.

[0045] According to a fourth aspect of the present invention, a computer program product is provided, the computer program product comprising a computer program, wherein when the computer program is executed, the above-described control method is performed.

[0046] In summary, the beneficial effects of the present invention are as follows:

[0047] This invention provides an active vibration control system and method for tiltrotor configurations. The active control system includes actuator components, an acceleration sensor, and an adaptive control unit. A piezoelectric laminate-based actuator component is used at the main wing spars to drive the spars structure and control bending moment and torque loading. A dual-loop control strategy and flight state tracking are employed. By coordinating rotor speed and rotor attitude with dual-loop control, the dynamic characteristics of the wing-rotor system and the steady-state response control of the airframe structure are achieved. This invention effectively improves the performance of the active vibration control system for tiltrotor configurations, expands its flight envelope applicability, and enhances flight safety and comfort of tiltrotor configurations. Attached Figure Description

[0048] Figure 1 This is a schematic diagram illustrating the composition of a preferred embodiment of the tiltrotor wing vibration active control system of the present invention;

[0049] The components include: a first wing actuator assembly 1; a second wing actuator assembly 2; an acceleration sensor 3 located inside the cockpit; an acceleration sensor 4 located at the outer end of the tiltrotor configuration wing; and an adaptive control unit 5.

[0050] Figure 2 This is a schematic diagram of the actuator assembly mounted on the tiltrotor configuration wing according to a preferred embodiment of the present invention;

[0051] Among them: piezoelectric laminated actuator 1-1 at the upper edge of the first wing front spars and piezoelectric laminated actuator 1-2 at the lower edge of the first wing front spars; piezoelectric laminated actuator 1-3 at the upper edge of the first wing rear spars and piezoelectric laminated actuator 1-4 at the lower edge of the first wing rear spars.

[0052] piezoelectric laminated actuator 2-1 on the upper edge of the second wing front spars and piezoelectric laminated actuator 2-2 on the lower edge of the second wing front spars; piezoelectric laminated actuator 2-3 on the upper edge of the second wing rear spars and piezoelectric laminated actuator 2-4 on the lower edge of the second wing rear spars.

[0053] Figure 3 This is a schematic diagram of the working principle of the control system according to a preferred embodiment of the present invention;

[0054] Figure 4 This is a schematic diagram of the composition of the adaptive control unit according to a preferred embodiment of the present invention;

[0055] Figure 5 This is a flowchart illustrating the selection of the accelerometer position according to a preferred embodiment of the present invention.

[0056] Figure 6 This is a schematic diagram illustrating the active vibration control effect of a preferred embodiment of the present invention. Detailed Implementation

[0057] To make the objectives, technical solutions, and advantages of this application clearer, the following detailed description is provided in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative and not intended to limit the scope of this application.

[0058] It should be noted that if the embodiments of the present invention involve directional indicators (such as up, down, left, right, front, back, etc.), the directional indicators are only used to explain the relative positional relationship and movement of the components in a certain specific posture (as shown in the figure). If the specific posture changes, the directional indicators will also change accordingly.

[0059] Furthermore, if the embodiments of this invention involve descriptions such as "first" or "second," these descriptions are for descriptive purposes only and should not be construed as indicating or implying their relative importance or implicitly specifying the number of technical features indicated. Therefore, a feature defined with "first" or "second" may explicitly or implicitly include at least one of those features. Additionally, the technical solutions of the various embodiments can be combined with each other, but this must be based on the ability of those skilled in the art to implement them. If the combination of technical solutions is contradictory or impossible to implement, it should be considered that such a combination of technical solutions does not exist and is not within the scope of protection claimed by this invention.

[0060] This invention provides an active vibration control system and method for a tiltrotor configuration. To make the above-mentioned objectives, features, and advantages of this invention more apparent and understandable, the invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. Please refer to... Figure 1 , Figure 2 , Figure 3 , Figure 4 .

[0061] The tilt rotor configuration vibration active control system is as follows: Figure 1 As shown, this includes actuator assemblies mounted on the wing spars, such as... Figure 2 As shown, the adaptive control unit is as follows: Figure 4 The figure shows an acceleration sensor measurement system. Figure 1 In the diagram, 1 and 2 are actuator components, 3 is the wing measuring point acceleration sensor, 4 is the fuselage measuring point acceleration sensor, and 5 is the adaptive control unit;

[0062] The active wing is implemented using a piezoelectric laminated actuator to drive the wing spars structure and its piezoelectric drive power supply, such as... Figure 2 As shown, it can realize the control of bending moment loading and torque loading of wing structure;

[0063] 201. The piezoelectric laminated actuator driving wing spars structure adopts a piezoelectric laminated actuator connected in parallel with the wing spars and installed at a certain angle with the elastic axis of the wing spars; the included angle is used to generate torque, to excite or suppress the torsional vibration mode of the wing, and to have a coupling control effect on bending vibration;

[0064] 202. The piezoelectric laminated actuator driving spar structure employs piezoelectric laminated actuators arranged in pairs above and below the spar's elastic axis, with each pair of actuators having an opposite angle to the elastic axis, such as... Figure 2 The first wing front spars upper edge piezoelectric laminated actuator 1-1 and the first wing rear spars upper edge piezoelectric laminated actuator 1-3, the first wing front spars lower edge piezoelectric laminated actuator 1-2 and the first wing rear spars lower edge piezoelectric laminated actuator 1-4, the second wing front spars upper edge piezoelectric laminated actuator 2-1 and the second wing rear spars upper edge piezoelectric laminated actuator 2-3, the second wing front spars lower edge piezoelectric laminated actuator 2-2 and the second wing rear spars lower edge piezoelectric laminated actuator 2-4;

[0065] 203. The piezoelectric laminated actuator driving spar structure employs piezoelectric laminated actuators arranged in pairs before and after the spar's elastic axis, with each pair of actuators having an opposite angle to the elastic axis, such as... Figure 2 The first wing front spars upper edge piezoelectric laminated actuator 1-1 and the first wing front spars lower edge piezoelectric laminated actuator 1-2, the first wing rear spars upper edge piezoelectric laminated actuator 1-3 and the first wing rear spars lower edge piezoelectric laminated actuator 1-4, the second wing front spars upper edge piezoelectric laminated actuator 2-1 and the second wing rear spars upper edge piezoelectric laminated actuator 2-3, the second wing rear spars upper edge piezoelectric laminated actuator 2-3 and the second wing rear spars lower edge piezoelectric laminated actuator 2-4;

[0066] 204. The bending moment loading control of the wing structure includes vertical bending moment loading control and yaw bending moment loading control;

[0067] 205. The vertical bending moment loading control of the wing structure is achieved by piezoelectric laminated actuators arranged in pairs aft and rear of the elastic shaft, such as... Figure 2 The piezoelectric multilayer actuator 1-1 on the upper edge of the first wing front spars and the piezoelectric multilayer actuator 1-2 on the lower edge of the first wing front spars, the piezoelectric multilayer actuator 1-3 on the upper edge of the first wing rear spars and the piezoelectric multilayer actuator 1-4 on the lower edge of the first wing rear spars, the piezoelectric multilayer actuator 2-1 on the upper edge of the second wing front spars and the piezoelectric multilayer actuator 2-3 on the upper edge of the second wing rear spars, and the piezoelectric multilayer actuator 2-3 on the upper edge of the second wing rear spars and the piezoelectric multilayer actuator 2-4 on the lower edge of the second wing rear spars are reverse driven to achieve this.

[0068] 206. The wing structure's directional bending moment loading control is achieved by piezoelectric laminated actuators arranged in pairs above and below the neutral axis, such as... Figure 2The piezoelectric multilayer actuator 1-1 on the upper edge of the first wing front spars and the piezoelectric multilayer actuator 1-3 on the upper edge of the first wing rear spars, the piezoelectric multilayer actuator 1-2 on the lower edge of the first wing front spars and the piezoelectric multilayer actuator 1-4 on the lower edge of the first wing rear spars, the piezoelectric multilayer actuator 2-1 on the upper edge of the second wing front spars and the piezoelectric multilayer actuator 2-3 on the upper edge of the second wing rear spars, and the piezoelectric multilayer actuator 2-2 on the lower edge of the second wing front spars and the piezoelectric multilayer actuator 2-4 on the lower edge of the second wing rear spars are reverse driven to achieve this.

[0069] 207. The torque loading control of the wing structure is achieved by piezoelectric laminated actuators arranged in pairs, front and rear, and top and bottom, along the neutral axis, such as... Figure 2 The piezoelectric actuator 1-1 at the upper edge of the first wing front spars and the piezoelectric actuator 1-4 at the lower edge of the first wing rear spars, the piezoelectric actuator 1-2 at the lower edge of the first wing front spars and the piezoelectric actuator 1-3 at the upper edge of the first wing rear spars, the piezoelectric actuator 2-1 at the upper edge of the second wing front spars and the piezoelectric actuator 2-4 at the lower edge of the second wing rear spars, and the piezoelectric actuator 2-2 at the lower edge of the second wing front spars and the piezoelectric actuator 2-3 at the upper edge of the second wing rear spars are driven in opposite directions.

[0070] 208. Under spatial constraints, the piezoelectric laminated actuator is arranged far from the wing neutral axis to maximize the driving efficiency of the piezoelectric drive unit and efficiently realize the suppression of dynamic load transmission through the wing and the response to wing bending and torsional vibrations.

[0071] like Figure 5 As shown, the acceleration sensor is placed at a key position on the fuselage to collect vibration acceleration signals of the fuselage and input them to the adaptive control unit to calculate control efficiency, control commands, etc.

[0072] Select initial candidate installation locations for the accelerometers, and based on factors such as cost and space, initially determine the number of accelerometers to be installed and combine them.

[0073] The accelerometer sensor should be positioned optimally. This involves initially selecting and filtering suitable sensor installation locations and measurement directions, with the following optimization rules:

[0074] The vibration control direction at the evaluation point considers three directions: X, Y, and Z.

[0075] A single accelerometer is unidirectional;

[0076] The symmetrical position sensors are a pair, and are selected simultaneously.

[0077] Using the above rules, the initial locations are screened, and 2n (n is a positive integer) combinations of accelerometer sensors are calculated to obtain the optimal location result.

[0078] The fuselage acceleration sensor is based on the finite element model, using the rotor hub force's primary channel vibration response at the evaluation point and the actuator unit's secondary channel response at the evaluation point as input conditions. Considering the above rules, a suitable combination of acceleration sensors is calculated.

[0079] 301. The adaptive control unit includes a piezoelectric drive power supply, an adaptive controller, and a signal conditioner; the signal conditioner has an AD unit for acquiring signals from an acceleration sensor and outputting them to the adaptive controller; the adaptive controller has a control signal output unit (including a DA unit) and a control algorithm calculation unit for providing the signal samples from the signal conditioner to the control algorithm, which calculates the control output and outputs the control signal through the DA unit; the piezoelectric drive power supply has a signal power amplification unit for receiving the control signal from the adaptive controller, amplifying it, and driving the actuator assembly to implement the output of control force to counteract structural vibration.

[0080] 302. The control strategy is a dual-loop control and flight status tracking strategy. The dual-loop control includes a first control loop and a second control loop. The first control loop implements dynamic characteristic control for the wing-rotor system, and the second control loop implements steady-state response control for the airframe structure. Flight status tracking includes rotor speed and attitude tracking.

[0081] 303. In the dual-loop control, the first control loop improves the convergence and control performance of the second control loop by controlling the dynamic characteristics of the wing-rotor system;

[0082] 304. The first control loop uses the modal response of the wing-rotor system as the feedback control variable to control the transient response characteristics of the tiltrotor configuration through the active wing;

[0083] 305. The first control loop is mainly for the low-order modes of the rotor-wing system, and can be selected as the first-order vertical bending mode, the first-order yaw bending mode, and the first-order torsional mode;

[0084] 306. The first control loop controls the pass frequency of the tilt rotor configuration non-rotor and the vibration response to transient excitation during tilt transition flight;

[0085] 307. The control algorithm of the first control loop adopts a time-domain feedback control algorithm, and optionally adopts a modal control or a linear quadratic Gaussian (LQG) control algorithm;

[0086] 308. The second control loop uses the steady-state response of the airframe structure as the feedback control variable, and controls the steady-state vibration response of the airframe through the active wing;

[0087] 309. The second control loop controls the steady-state harmonic vibration response of the tilt rotor configuration rotor through frequencies and their integer multiples;

[0088] 310. The second control loop control algorithm adopts a recursive frequency domain optimal feed control algorithm. U k and U k-1 Here, μ represents the control input harmonic coefficients for the k-th control correction time step and the previous step, and μ is the correction step size. Let be the optimal control harmonic coefficient vector at time k, where The transfer matrix from the piezoelectric drive control voltage of the main engine wing to the acceleration response at the vibration observation point is... These are the control point response weighting matrix and the control input weighting matrix, respectively. The vector of harmonic coefficients of the acceleration response at the control point at time k-1;

[0089] 311. The rotor speed and attitude tracking are provided by the rotor speed sensor signal and the rotor tilt controller in the flight control system, such as... Figure 3 The external input signal is shown in the figure;

[0090] 312. The rotor speed tracking is used for switching the control model and generating reference signals in the second control loop;

[0091] 313. The rotor attitude tracking is used for coordination between the first control loop and the second control loop;

[0092] 314. Coordination between the first control loop and the second control loop: in helicopter mode, the first control loop and the second control loop have the same weight; in transition mode, the weight of the first control loop is increased; and in airplane mode, the weight of the second control loop is increased.

[0093] The sensing and measurement system includes a vibration acceleration sensor, a signal conditioner, and necessary flight control system signals;

[0094] 401. The acceleration sensor includes an acceleration sensor placed on the body (such as...). Figure 1 (as shown in Figure 3) and acceleration sensors placed on the wings (such as...) Figure 1 As shown in Figure 4, the acceleration sensor signal placed on the fuselage is used for the second control loop, and the acceleration sensor signal placed on the wing is used for both the first and second control loops.

[0095] 402. The arrangement of the wing acceleration sensors is based on the wing modal characteristics of the selected control mode, and is arranged at the peak of the selected control mode. It can be selected according to the principle of importance of low-order modes of vibration control. Generally, three sensors are arranged at the wingtip: two vertical and one chordal. The two vertical sensors are arranged on the front and rear wing spars respectively, for measuring the vertical bending, chord bending and torsional vibration response of the wing. The fuselage acceleration sensors are arranged in the key instruments in the cockpit, the pilot's position (including the foot pedals), and the cockpit seat position. The number of sensors is controlled between 8 and 15 to take into account the distribution rate of the control points of interest and the complexity of the control system.

[0096] 403. The necessary flight control system signals are the rotor speed sensor signals and nacelle tilt control signals in the flight control system.

[0097] According to a second aspect of the present invention, a method for active control of tiltrotor wing vibration is proposed, employing the aforementioned active control system for tiltrotor wing vibration, comprising the following steps:

[0098] Step 1: Signal Acquisition and Conditioning. The wing acceleration sensor and fuselage acceleration sensor acquire vibration acceleration signals in real time. After being amplified and filtered by the signal conditioner, the signals are input to the adaptive controller. At the same time, the adaptive controller obtains rotor speed signals and nacelle tilt attitude signals from the flight control system.

[0099] Step 2: Flight Status Determination and Control Strategy Decision. The adaptive controller determines the current flight mode based on the acquired rotor speed and nacelle tilt angle, and determines the weight allocation of the dual-loop control strategy; wherein, the dual-loop control strategy includes a first control loop and a second control loop, the first control loop implements dynamic characteristic control for the wing-rotor system, and the second control loop implements steady-state response control for the airframe structure;

[0100] Step 3: Obtain the time-domain vibration signal. The adaptive controller samples the conditioned acceleration signal to obtain the time-domain vibration acceleration signal e(k) at the current sampling time k;

[0101] Step 4: Execute the first control loop (dynamic characteristic control). Based on the flight state determination result, execute the first control loop, using the signal from the wing acceleration sensor as input, and employing a time-domain feedback control algorithm to calculate the wing modal control quantity. The control object of the first control loop mainly targets the low-order modes of the rotor-wing system, including the first-order vertical bending mode, the first-order yaw bending mode, and the first-order torsional mode, used to suppress the vibration response of the non-rotor passing frequency and the transient excitation during tilt transition flight.

[0102] Step 5: Execute the second control loop (steady-state response control). Based on the flight state determination result, execute the second control loop, using the signal from the fuselage acceleration sensor as input, and employing a recursive frequency domain optimal feedforward control algorithm to calculate the steady-state harmonic control quantity;

[0103] Step Six: Control Signal Synthesis and Output. The control quantities of the first control loop obtained in Step Four and the second control loop obtained in Step Five are weighted and fused according to the weights determined in Step Two to obtain the total control input harmonic vector; the time-domain control output is calculated based on the total control input harmonic vector, and the analog control signal is output through the DA unit.

[0104] Step 7: Power Amplification and Actuation. The control signal is amplified by the piezoelectric drive power supply and applied to the piezoelectric multilayer actuator, causing it to output an active force with the opposite phase to the control target, thereby implementing vibration control;

[0105] Step 8: Iterative Loop Control. Measure the vibration response after control in real time using an accelerometer, and repeat steps 1 through 7 to form a closed-loop control until the vibration converges within the target range.

[0106] In one possible embodiment, step two, the flight state determination and control strategy decision, specifically includes:

[0107] a) Determine the flight mode based on the nacelle tilt angle θ: when θ≥75°, it is determined to be helicopter mode; when 75°>θ>15°, it is determined to be tilt transition mode; when θ≤15°, it is determined to be airplane mode.

[0108] b) Determine the weights of the dual-loop control strategy based on the flight mode: In helicopter mode, the first and second control loops have the same weight (W1=W2=0.5); in tilt transition mode, the weight of the first control loop increases (W1=0.7-0.8, W2=0.2-0.3); in airplane mode, the weight of the second control loop increases (W1=0.2-0.3, W2=0.7-0.8).

[0109] c) Determine the target control frequency of the second control loop based on the rotor speed signal Ω, i.e. the rotor passing frequency 1Ω and its integer multiples 2Ω, 3Ω, ..., NΩ, where N is a positive integer;

[0110] d) Based on the nacelle tilt angle signal, switch the weight of the acceleration sensor at different positions, and increase the weight of the wing acceleration sensor in the control target in the tilt transition mode to enhance the suppression of transient response.

[0111] In one possible embodiment, in step four, the time-domain feedback control algorithm used in the first control loop is either a linear quadratic Gaussian (LQG) control algorithm or a modal control algorithm. Using the wing modal coordinates η(t) as the state variable, a state-space equation is constructed. The system state is estimated using a Kalman filter, and the optimal control force is calculated based on the linear quadratic regulator (LQR) to achieve active damping injection of the selected low-order modes, thereby altering the transient response characteristics of the wing-rotor system.

[0112] In one possible embodiment, in step four, the control object of the first control loop is preferentially selected as the low-order mode that has the greatest impact on flight safety and vibration transmission, including the first-order vertical bending mode, the first-order yaw bending mode, and the first-order torsional mode; the control priority of each mode is dynamically adjusted according to the flight mode, for example, in helicopter mode, the focus is on controlling the torsional mode, and in airplane mode, the focus is on controlling the vertical bending mode.

[0113] In one possible embodiment, in step five, the second control loop uses a recursive frequency domain optimal feedforward control algorithm, whose control frequency is updated in real time based on the rotor speed signal; when the rotor speed changes, the algorithm automatically adjusts the frequency components in the reference signal Φ(k) to achieve steady-state harmonic vibration tracking and suppression under variable speed conditions.

[0114] In one possible embodiment, in step six, the fusion of the first control loop and the second control loop adopts a weighted superposition method: the total control quantity u(t) = W1u1(t) + W2u2(t), where u1(t) is the time-domain control quantity output by the first control loop and u2(t) is the time-domain control quantity output by the second control loop; W1 and W2 satisfy W1 + W2 = 1 and are adjusted in real time according to the flight mode.

[0115] In one possible embodiment, step five, which involves calculating the steady-state harmonic control quantity using a recursive frequency domain optimal feedforward control algorithm, includes the following steps:

[0116] The time-domain vibration acceleration signal e(k) is given, where k is the current sampling time.

[0117] Update the control input harmonic coefficient vector using the sampled values ​​of the vibration acceleration signal at the current sampling time. The update formula used for calculation is as follows:

[0118]

[0119]

[0120]

[0121] In the formula, Let k be the acceleration response coefficient vector of the control point at time k. Let K be the acceleration response coefficient vector of the control point at time k-1. , To control the harmonic frequencies, i = 1, 2, ..., R, Forgetting factor, optional =0.98; Initial gain matrix It can be any positive definite matrix, and can be selected. =I is the identity matrix;

[0122] From the formula Calculate the optimal control input harmonic coefficients, and then update the control input harmonic vector. ,in The transfer matrix from the piezoelectric drive control voltage of the main engine wing to the acceleration response at the vibration observation point is... These are the control point response weighting matrix and the control input weighting matrix, respectively. Let μ be the harmonic coefficient vector of the acceleration response at the control point at time k-1, and μ be the correction step size. and The control input harmonic coefficients for the k-th control correction time step and the previous step;

[0123] Calculate time-domain control output The DA unit outputs control commands.

[0124] like Figure 6 As shown, the control signal is applied to the piezoelectric stacked actuator via the piezoelectric drive power supply to implement vibration control. Control is implemented from the 5th second. After calculation and convergence, a low vibration control effect is achieved. Specifically, the light-colored line part belongs to the single control loop method, which has a larger coverage area, indicating a long convergence process; the dark-colored line part belongs to the double control loop method, which has a smaller coverage area, indicating a fast convergence process.

[0125] The technical features of the above embodiments can be combined in any way. For the sake of brevity, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0126] The embodiments described above are merely illustrative of several implementation methods of this application, and while the descriptions are relatively specific and detailed, they should not be construed as limiting the scope of the invention patent. It should be noted that those skilled in the art can make various modifications and improvements without departing from the concept of this application, and these all fall within the protection scope of this application. Therefore, the protection scope of this patent application should be determined by the appended claims.

Claims

1. An active control system for the vibration of a tiltrotor wing, characterized in that, include: The system comprises an actuator assembly, acceleration sensors, and an adaptive control unit. The actuator assembly includes a plurality of piezoelectric laminated actuators, centrally symmetrically arranged within the tiltrotor configuration wings on both sides of the fuselage. The acceleration sensors include wing acceleration sensors located at the outer ends of the tiltrotor configuration wings and fuselage acceleration sensors located inside the cockpit, used to collect vibration source signals from the tiltrotor configuration wings and vibration responses from evaluation points within the cockpit, respectively. The adaptive control unit collects and conditions the signals from the acceleration sensors, outputs control commands to an internal piezoelectric drive power supply via an adaptive control algorithm, and outputs control energy to the actuator assembly via the piezoelectric drive power supply. The adaptive control algorithm employs a dual-loop control strategy: the first control loop uses a time-domain feedback control algorithm to implement dynamic characteristic control of the wing-rotor system, used to suppress non-rotor pass-through frequency and tilt transition flight. The transient excitation vibration response is assessed. The second control loop employs a recursive frequency domain optimal feedforward control algorithm to implement steady-state response control for the airframe structure, suppressing steady-state harmonic vibrations at the rotor pass frequency and its integer multiples. The adaptive control unit determines the flight mode based on the nacelle tilt angle θ: when θ ≥ 75°, it is determined to be helicopter mode; when 75° > θ > 15°, it is determined to be tilt transition mode; when θ ≤ 15°, it is determined to be aircraft mode. The weights of the dual-loop control strategy are determined based on the flight mode: in helicopter mode, the first and second control loops have the same weight; in tilt transition mode, the weight of the first control loop is increased; in aircraft mode, the weight of the second control loop is increased. The target control frequency of the second control loop is determined based on the rotor speed signal Ω, and the weights of acceleration sensors at different positions are switched based on the nacelle tilt angle signal. In tilt transition mode, the weight of the wing acceleration sensor in the control target is increased to enhance the suppression of transient response.

2. The active vibration control system for a tiltrotor wing according to claim 1, characterized in that, The tiltrotor configuration wing has a wing spars structure inside, which includes a front spars and a rear spars arranged parallel to each other along the span of the fuselage; multiple stringers are distributed parallel to each other between the front spars and the rear spars; the piezoelectric laminated actuators are arranged in pairs between adjacent stringers on the upper and lower edges of the front spars and the upper and lower edges of the rear spars.

3. The active vibration control system for a tiltrotor wing according to claim 2, characterized in that, The piezoelectric laminated actuator is connected in parallel with the front and rear beams.

4. The active vibration control system for a tiltrotor wing according to claim 2, characterized in that, The axial direction of the piezoelectric laminated actuator is at an angle to the front or rear beam. This angle is used to generate a torque, to excite or suppress the torsional vibration mode of the wing, and to have a coupling control effect on bending vibration. The axial direction of each pair of piezoelectric laminated actuators at the upper and lower edges of the front or rear beams is opposite to the angle between the front and rear beams. The axial direction of each pair of piezoelectric laminated actuators at the upper and lower edges of the front and rear beams, and at the lower and lower edges of the front and rear beams, is opposite to the angle between the front and rear beams.

5. The active vibration control system for a tiltrotor wing according to claim 1, characterized in that, The fuselage acceleration sensor is located at various locations, including critical instruments in the cockpit, the pilot's position, and the cockpit seat.

6. The active vibration control system for a tiltrotor wing according to claim 1, characterized in that, The wing acceleration sensor includes at least two vertical acceleration sensors and one chordal acceleration sensor. The two vertical acceleration sensors are respectively arranged on the front and rear spars of the wing and are used to measure the vertical bending, chordal bending and torsional vibration response of the wing. In order to effectively measure the chordal bending vibration of the wing, the chordal acceleration sensor is arranged in the wingtip region. The specific installation position includes, but is not limited to, the leading edge of the wingtip, the trailing edge of the wingtip, or the corresponding position of the winglet, and is used to sense and measure the vibration acceleration of the wing along the chord.

7. The active vibration control system for a tiltrotor wing according to claim 1, characterized in that, The adaptive control unit includes a piezoelectric drive power supply, an adaptive controller, and a signal conditioner. The signal conditioner has an AD unit for acquiring signals from the accelerometer and outputting them to the adaptive controller. The adaptive controller has a DA unit, a control signal output unit, and a control algorithm calculation unit for providing the signal samples from the signal conditioner to the control algorithm. The control algorithm program calculates the control output and outputs the control signal through the DA unit. The piezoelectric drive power supply has a signal power amplification unit for receiving the control signal from the adaptive controller, amplifying it, and then driving the actuator assembly to output control force and counteract structural vibration.

8. A method for active control of tiltrotor wing vibration, employing the active control system for tiltrotor wing vibration as described in any one of claims 1-7, characterized in that, Includes the following steps: Step 1: Signal Acquisition and Conditioning; The wing acceleration sensor and fuselage acceleration sensor respectively acquire vibration acceleration signals in real time, which are then amplified and filtered by the signal conditioner and input to the adaptive controller; At the same time, the adaptive controller obtains rotor speed signals and nacelle tilt attitude signals from the flight control system; Step Two: Flight State Determination and Control Strategy Decision; The adaptive controller determines the current flight mode based on the acquired rotor speed and nacelle tilt angle, and determines the weight allocation of the dual-loop control strategy; wherein, the dual-loop control strategy includes a first control loop and a second control loop, the first control loop implements dynamic characteristic control for the wing-rotor system, and the second control loop implements steady-state response control for the airframe structure; Step Two, specifically, flight state determination and control strategy decision includes: Flight mode is determined based on nacelle tilt angle θ: when θ ≥ 75°, it is determined to be helicopter mode; when 75° > θ > 15°, it is determined to be tilt transition mode; when θ ≤ 15°, it is determined to be airplane mode. The weights of the dual-loop control strategy are determined based on the flight mode: in helicopter mode, the first and second control loops have the same weight; in tilt transition mode, the weight of the first control loop increases; and in airplane mode, the weight of the second control loop increases. The target control frequency of the second control loop is determined based on the rotor speed signal Ω, i.e. the rotor passing frequency 1Ω and its integer multiples 2Ω, 3Ω, ..., NΩ, where N is a positive integer; The weights of acceleration sensors at different positions are switched according to the nacelle tilt angle signal. In the tilt transition mode, the weight of the wing acceleration sensor in the control target is increased to enhance the suppression of transient response. Step 3: Obtain the time-domain vibration signal; the adaptive controller samples the conditioned acceleration signal to obtain the time-domain vibration acceleration signal e(k) at the current sampling time k; Step 4: Based on the flight state determination result, execute the first control loop, using the signal from the wing acceleration sensor as input, and employing a time-domain feedback control algorithm to calculate the wing modal control quantity; the control object of the first control loop is the low-order modes of the rotor-wing system, including the first-order vertical bending mode, the first-order yaw bending mode, and the first-order torsional mode, used to suppress the vibration response of the non-rotor passing frequency and the transient excitation during tilt transition flight; Step 5: Based on the flight status judgment result, execute the second control loop, using the signal from the fuselage acceleration sensor as input, and calculate the steady-state harmonic control quantity using the recursive frequency domain optimal feedforward control algorithm; Step Six: Control Signal Synthesis and Output; The control quantities of the first control loop obtained in Step Four and the second control loop obtained in Step Five are weighted and fused according to the weights determined in Step Two to obtain the total control input harmonic vector; The time-domain control output is calculated based on the total control input harmonic vector, and the analog control signal is output through the DA unit; Step 7: Power Amplification and Actuation; The control signal is amplified by the piezoelectric drive power supply and applied to the piezoelectric stacked actuator, causing it to output an active force with the opposite phase to the control target, thereby implementing vibration control; Step 8: Iterative loop; measure the vibration response after control in real time using an accelerometer, repeat steps 1 to 7 to form a closed-loop control until the vibration converges to the target range.

9. The method according to claim 8, characterized in that: In step four, the time-domain feedback control algorithm used in the first control loop is either a linear quadratic Gaussian control algorithm or a modal control algorithm.

10. The method according to claim 8, characterized in that: In step four, the control objects of the first control loop are preferentially selected to be the low-order modes that have the greatest impact on flight safety and vibration transmission, including the first-order vertical bending mode, the first-order yaw bending mode, and the first-order torsional mode. The control priority of each mode is dynamically adjusted according to the flight mode. In helicopter mode, the focus is on controlling the torsional mode, and in airplane mode, the focus is on controlling the vertical bending mode.

11. The method according to claim 8, characterized in that: In step five, the second control loop uses a recursive frequency domain optimal feedforward control algorithm, whose control frequency is updated in real time based on the rotor speed signal. When the rotor speed changes, the algorithm automatically adjusts the frequency components in the reference signal Φ(k) to achieve steady-state harmonic vibration tracking and suppression under variable speed conditions.

12. The method according to claim 8, characterized in that: In step six, the fusion of the first control loop and the second control loop adopts a weighted superposition method: the total control quantity u(t) = W1u1(t) + W2u2(t), where u1(t) is the time-domain control quantity output by the first control loop and u2(t) is the time-domain control quantity output by the second control loop; W1 and W2 satisfy W1 + W2 = 1 and are adjusted in real time according to the flight mode.

13. The method according to claim 8, characterized in that: In step five, the specific process of calculating the steady-state harmonic control quantity using the recursive frequency domain optimal feedforward control algorithm includes: The time-domain vibration acceleration signal e(k) is given by k, where k is the current sampling time. Update the control input harmonic coefficient vector using the sampled values ​​of the vibration acceleration signal at the current sampling time. The update formula used for calculation is as follows: , , ; In the formula, Let k be the acceleration response coefficient vector of the control point at time k. Let K be the acceleration response coefficient vector of the control point at time k-1. , To control the harmonic frequencies, i = 1, 2, ..., R, Forgetting factor, initial gain matrix It can be any positive definite matrix, and can be selected. =I is the identity matrix; From the formula Calculate the optimal control input harmonic coefficients, and then update the control input harmonic vector. ,in The transfer matrix from the piezoelectric drive control voltage of the main engine wing to the acceleration response at the vibration observation point is... These are the control point response weighting matrix and the control input weighting matrix, respectively. Let μ be the harmonic coefficient vector of the acceleration response at the control point at time k-1, and μ be the correction step size. and The control input harmonic coefficients for the k-th control correction time step and the previous step; Calculate time-domain control output The DA unit outputs control commands.

14. A computer-readable storage medium, characterized in that, The computer-readable storage medium is used to store a computer program, which, when executed, performs the method according to any one of claims 8-13.

15. A computer program product, characterized in that, The computer program product includes a computer program that, when executed, performs the method according to any one of claims 8-13.

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