Active control method for aircraft engine rotor maneuvering flight

Abstract

The present invention relates to the technical field of rotor vibration control, and more specifically to a method for actively controlling aircraft engine rotor maneuvering flight, comprising: obtaining system attribute parameters and operating state parameters of the aircraft engine; constructing an aircraft engine rotor maneuvering flight dynamics model with a master-controlled elastic support friction damper, and obtaining the aircraft engine rotor vibration amplitude; obtaining the rotor vibration amplitude error at each moment; obtaining the operating conditions of the maneuvering flight; and obtaining a corresponding pressure control signal based on the operating conditions of the maneuvering flight and controlling the aircraft engine damper. The present invention provides targeted and responsive precise control of rotor vibration.

Description

Technical Field

[0001] The present invention relates to the technical field of rotor vibration control, and in particular to an active control method for maneuvering flight of an aero-engine rotor. Background Art

[0002] Aircraft engines are a comprehensive reflection of a country's industrial foundation, scientific and technological level, and overall national strength, and therefore hold a crucial position. To meet the high-performance and high-reliability demands of modern engines, engine rotor systems are increasingly evolving toward high-speed, lightweight structures. Modern high-speed aircraft engines commonly utilize rigid rotors mounted on flexible supports—twin rotor engines with intermediate supports. These engines feature complex structures with intermediate bearings and multiple stages of disk blades (turbine, fan, and compressor blades). These engines must withstand complex operating environments such as high speeds, dual-frequency excitation (high and low pressure), strong coupling between high and low pressure rotors, aerodynamic and thermal stresses, and the unique operating conditions of maneuverable flight.

[0003] With the application and pre-research of fourth- and fifth-generation aircraft, aircraft maneuverability has become a key indicator in the evaluation system of new fighter jets. Typical aircraft maneuvers include level flight acceleration, jump maneuvers, circling maneuvers, and somersault maneuvers. The high maneuverability of aircraft poses severe challenges to the design, operation, and maintenance of aircraft engines. In maneuverable flight environments, aircraft engines not only experience high speeds and accelerations, but also high maneuvering accelerations. Therefore, when the aircraft completes various maneuvers, the engine rotor is not only subjected to various excitations of its own, but also to additional excitation forces, namely, additional maneuvering inertia load excitations. These effects can cause a previously stable rotor system to suddenly experience large vibrations, which can lead to a series of failures such as rotor instability, rubbing of moving and static parts, and blade breakage.

[0004] To avoid the above phenomenon, aircraft engine maneuvering flight requires a certain vibration control capability. The existing technology has some vibration control of vibration reduction and load reduction mechanisms, but they are mostly passive control and cannot fully meet the vibration control needs of all rotor modal maneuvering flight. For traditional linear control methods, such as PID (proportional-integral-differential control), H ∞ , LQR (linear quadratic control) methods will find it difficult to meet the needs of rapid response, and traditional nonlinear control methods such as intelligent control methods are difficult to achieve ideal results in actual work due to their complex structure and large amount of calculation.

[0005] Therefore, it is necessary to provide an active control method for aircraft engine rotor maneuvering flight to solve the above problems. Summary of the Invention

[0006] The present invention provides an active control method for aircraft engine rotor maneuvering flight to solve the problems of traditional linear control methods such as PID (proportional-integral-differential control), H ∞ , LQR (linear quadratic control) methods will find it difficult to meet the needs of rapid response, and traditional nonlinear control methods such as intelligent control methods are difficult to achieve ideal results in actual work due to their complex structure and large amount of calculation.

[0007] The present invention provides an active control method for aircraft engine rotor maneuvering flight, which adopts the following technical solutions, including:

[0008] Obtaining system attribute parameters and operating state parameters of the aircraft engine; the operating state parameters include: the common working line of the engine rotor, i.e., the rotation speed of each operating condition, and the related maneuvering flight action types, wherein the maneuvering flight action types include: turning and circling, diving and pulling up;

[0009] Based on the system attribute parameters and operating state parameters of the aircraft engine, a flight dynamics model of the aircraft engine rotor maneuver with a main control type elastic support dry friction damper is established; based on the flight dynamics model of the aircraft engine rotor maneuver with a main control type elastic support dry friction damper, the rotor vibration amplitude of the aircraft engine is obtained;

[0010] The amplitude extraction method is used to separate the single-frequency amplitude component, double-frequency amplitude component, and triple-frequency amplitude component of the rotor vibration amplitude; the rotor vibration amplitude error at each moment is obtained based on the maximum amplitude component among the single-frequency amplitude component, double-frequency amplitude component, and triple-frequency amplitude component and the reference rotor amplitude;

[0011] Obtain the overload factor of the aircraft during maneuvering flight based on the roll angle, and obtain the operating conditions of the maneuvering flight based on the overload factor and the preset overload factor threshold. The operating conditions include normal operating conditions, non-emergency operating conditions, and emergency operating conditions;

[0012] When the operating condition of the maneuvering flight is a normal operating condition, the damper of the aircraft engine is controlled according to the set positive pressure control signal; when the operating condition of the maneuvering flight is a non-emergency operating condition, the pressure control signal is output according to the rotor vibration amplitude error and the PI control method, and the damper of the aircraft engine is controlled according to the pressure control signal; when the operating condition of the maneuvering flight is an emergency operating condition, the pressure control signal is output according to the rotor vibration amplitude error and the MFAC control method, and the damper of the aircraft engine is controlled according to the pressure control signal.

[0013] Preferably, the system attribute parameters of the aircraft engine include: a system stiffness matrix, a system damping matrix, a system mass matrix, and a system gyroscopic torque matrix.

[0014] Preferably, establishing a maneuvering flight dynamics model of an aeroengine rotor with a master-controlled elastic support friction damper includes:

[0015] A rotor maneuvering flight model of an aircraft engine is constructed based on the rotor speed, the angular velocity component in the x-axis direction, the system attribute parameters of the aircraft engine, and the rotor vibration amplitude, rotor vibration velocity, and rotor vibration acceleration.

[0016] The main control type elastic support dry friction damper model is adopted as the main control type elastic support dry friction damper dynamic model;

[0017] According to the main control type elastic support dry friction damper model and the rotor maneuvering flight model, the aircraft engine rotor maneuvering flight dynamics model with the main control type elastic support dry friction damper is constructed.

[0018] Preferably, the rotor maneuver flight model of the aircraft engine is expressed as:

[0019]

[0020] Where, is the rotor speed, is the angular velocity component of the rotor in the horizontal x direction, which is affected by the maneuvering flight action. is the system mass matrix; is the system damping matrix; is the system stiffness matrix; is the system gyro torque matrix; is the parameter stiffness matrix related to the rotor speed and the angular velocity component of the rotor angular velocity in the horizontal x direction; For external forces; is the rotor vibration amplitude; is the rotor vibration speed; is the rotor vibration acceleration.

[0021] Preferably, obtaining the master-controlled elastic support friction damper model includes:

[0022] The expression of the main control type elastic support dry friction damper model in the viscous state is:

[0023]

[0024] The expression of the main control type elastic dry friction damper model in the slip state is:

[0025]

[0026] Where, represents the equivalent stiffness of the damper in the viscous state; It represents the equivalent stiffness of the damper in the slip state; represents the equivalent damping of the damper in the viscous state; It represents the equivalent damping of the damper in the sliding state; Indicates the absolute value of the response of the dynamic friction plate in the aircraft engine system; Indicates the contact stiffness of the static friction plate and the dynamic friction plate; Indicates the installation stiffness of the static friction plate; Indicates the mass of the static friction plate; represents the coefficient of friction; Indicates positive pressure.

[0027] Preferably, the expression of the flight dynamics model of the rotor maneuver of an aero-engine with a main-controlled elastic dry friction damper is:

[0028]

[0029] Where, is the rotor speed, is the angular velocity component of the rotor in the horizontal x direction, which is affected by the maneuvering flight action. is the system mass matrix; is the system damping matrix; is the system stiffness matrix; is the system gyro torque matrix; is the parameter stiffness matrix related to the rotor speed and the angular velocity component of the rotor angular velocity in the horizontal x direction; For external forces; is the rotor vibration amplitude; is the rotor vibration speed; is the rotor vibration acceleration; is the equivalent stiffness of the damper, Pick or , is the equivalent stiffness of the damper; is the equivalent damping of the damper, Pick or .

[0030] Preferably, in non-emergency conditions, the pressure control signal is expressed as:

[0031]

[0032] Where, is the pressure control signal at the nth moment in non-emergency conditions; is the proportional coefficient of PI control; is the integral coefficient of PI control; is the rotor vibration amplitude error at the nth moment.

[0033] Preferably, in an emergency condition, the pressure control signal is expressed as:

[0034]

[0035] Where, is the pressure control signal at the nth moment in the emergency condition; is the partial derivative; is the weight coefficient; is the step length; is the rotor vibration amplitude error between the ideal vibration amplitude and the actual vibration amplitude at the n+1th moment; is the number of iterations.

[0036] Preferably, the step of obtaining the maneuvering flight condition according to the overload coefficient and a preset overload coefficient threshold range is:

[0037] The preset overload coefficient thresholds include: a first overload coefficient threshold, a second overload coefficient threshold, and a third overload coefficient threshold, and the first overload coefficient threshold is smaller than the second overload coefficient threshold and smaller than the third overload coefficient threshold;

[0038] When the overload factor is greater than or equal to the first overload factor threshold and less than the second overload factor threshold, the aircraft engine is in a normal operating condition;

[0039] When the overload factor is greater than or equal to the second overload factor threshold and less than the third overload factor threshold, the aircraft engine is in a non-emergency operating condition;

[0040] When the overload factor is greater than or equal to the third overload factor threshold, the aircraft engine is in an emergency operating condition.

[0041] Preferably, the RLS algorithm is used to separate the single-frequency amplitude component, the double-frequency amplitude component and the triple-frequency amplitude component of the rotor vibration amplitude.

[0042] The beneficial effects of the present invention are:

[0043] In view of the fact that most of the reference positions in the existing ESDFD control method are fixed, the rotor mode is not considered, and the vibration level of the rotor cannot be reflected, the present invention applies a main-controlled elastic dry friction damper to the rotor maneuvering flight, and combines it with the aircraft engine rotor maneuvering flight model to obtain an aircraft engine rotor maneuvering flight model with a main-controlled elastic dry friction damper. The aircraft engine rotor maneuvering flight model with the main-controlled elastic dry friction damper is used to obtain the rotor vibration amplitude, and the rotor vibration amplitude error is obtained based on the various frequency-harmonic amplitude components of the rotor vibration amplitude and the rotor reference vibration amplitude. Then, the maneuvering flight conditions are divided. When the maneuvering flight conditions are normal conditions, the damper of the aircraft engine is controlled according to the set positive pressure control signal; when the maneuvering flight conditions are non-emergency conditions, a pressure control signal is output according to the rotor vibration amplitude error and the PI control method, and the damper of the aircraft engine is controlled according to the pressure control signal. When the maneuvering flight conditions are emergency conditions, the control system needs to respond quickly, and traditional linear control methods such as PID (proportional-integral-differential control) and H ∞ The LQR (Linear Quadratic Control) method requires multiple adjustments and is difficult to meet the requirements of rapid response. Furthermore, traditional nonlinear control methods, such as intelligent control methods, are difficult to achieve ideal results in practice due to their complex structures and large computational workload. MFAC, on the other hand, is essentially a data-driven algorithm with strong engineering practicality. Therefore, this paper outputs a pressure control signal based on the rotor vibration amplitude error and the MFAC control method. This pressure control signal controls the aircraft engine's damper, thereby achieving targeted and responsive precise control of rotor vibration. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0045] Figure 1 This is a flow chart of a method for actively controlling rotor maneuvering flight of an aero-engine according to the present invention;

[0046] Figure 2 This is a control block diagram of an active control method for aircraft engine rotor maneuvering flight according to the present invention;

[0047] Figure 3 This is a force analysis diagram of the aircraft turning and circling of the present invention;

[0048] Figure 4This is a force analysis diagram of the aircraft of the present invention when diving or pulling up;

[0049] Figure 5 This is a flowchart of the amplitude extraction method based on the RLS algorithm;

[0050] Figure 6 is a relationship diagram between the overload factor and the roll angle of the aircraft under normal operating conditions of the present invention;

[0051] Figure 7 It is a structural schematic diagram of the master-controlled elastic support dry friction damper of the present invention.

[0052] In the figure: 1. Dynamic friction plate; 2. Sealing ring; 3. Mounting ring; 4. Static friction plate; 5. Piezoelectric ceramic actuator; 6. Damper mounting cylinder; 7. Engine load-bearing frame; 8. Engine elastic support; 9. Engine bearing. DETAILED DESCRIPTION

[0053] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0054] An embodiment of an active control method for aircraft engine rotor maneuvering flight of the present invention is as follows: Figure 1 As shown, including:

[0055] S1. Obtaining system attribute parameters and operating status parameters of the aircraft engine;

[0056] Specifically, the system attribute parameters of the aircraft engine and the operating status parameters of the aircraft engine are obtained; the operating status parameters include: the common working line of the engine rotor, that is, the rotation speed of each operating condition and the related maneuvering flight action type, among which the maneuvering flight action type includes: turning and circling, diving and pulling up.

[0057] Among them, the system attribute parameters of the aircraft engine include: system stiffness matrix, system damping matrix, system mass matrix, and system gyroscopic torque matrix.

[0058] like Figure 3 and Figure 4 As shown in the figure, common maneuvering flight actions include turning and circling, diving and pulling up. The force analysis diagram of turning and circling is shown in the figure. Figure 3 As shown in the figure, the force analysis diagram of the aircraft when diving or pulling up is as follows Figure 4 shown.

[0059] S2. Construct a maneuvering flight dynamics model of an aeroengine rotor with a master-controlled elastic support friction damper and obtain the rotor vibration amplitude of the aeroengine;

[0060] Specifically, a rotor maneuvering flight dynamics model of an aero-engine with a main-controlled elastic-braced dry friction damper is established according to the system property parameters and the operating state parameters of the aero-engine. The rotor vibration amplitude of the aero-engine is obtained according to the rotor maneuvering flight dynamics model of an aero-engine with a main-controlled elastic-braced dry friction damper.

[0061] Step 21, the steps of establishing the rotor maneuvering flight dynamics model of an aircraft engine with a master-controlled elastic-branched friction damper are as follows: constructing the rotor maneuvering flight model of the aircraft engine according to the rotor speed of the aircraft engine, the angular velocity component in the x-axis direction, the system attribute parameters of the aircraft engine, and the rotor vibration amplitude, rotor vibration velocity and rotor vibration acceleration; using the master-controlled elastic-branched friction damper model as the master-controlled elastic-branched friction damper dynamics model; constructing the rotor maneuvering flight dynamics model of the aircraft engine with the master-controlled elastic-branched friction damper according to the master-controlled elastic-branched friction damper model and the rotor maneuvering flight model.

[0062] The differential equations of motion of the rotor system under different maneuvering flight states considered in this embodiment are:

[0063] (1)

[0064] Where, is the rotor speed, is the angular velocity component of the rotor in the horizontal x direction, which is affected by the maneuvering flight action. is the system mass matrix; is the system damping matrix; is the system stiffness matrix; is the system gyro torque matrix; is the parameter stiffness matrix related to the rotor speed and the angular velocity component of the rotor angular velocity in the horizontal x direction; For external forces; is the rotor vibration amplitude; is the rotor vibration speed; is the rotor vibration acceleration.

[0065] According to the rotor vibration amplitude , vibration speed , vibration acceleration is The initial value at time zero is used to calculate the transient vibration response of the rotor under maneuvering flight conditions using the Newmark-HHT method. For each time step, the numerical integration requires recalculating the unbalanced force. , basic exciting force , oil film force And there is also the external force caused by the work done by gravity The rotor system motion differential equation is used as the rotor maneuver flight model of the aircraft engine.

[0066] The acquisition of the master-controlled elastic-supported dry friction damper model includes:

[0067] The expression of the main control type elastic support dry friction damper model in the viscous state is:

[0068] (2)

[0069] The expression of the main control type elastic dry friction damper model in the slip state is:

[0070] (3)

[0071] Where, represents the equivalent stiffness of the damper in the viscous state; It represents the equivalent stiffness of the damper in the slip state; represents the equivalent damping of the damper in the viscous state; It represents the equivalent damping of the damper in the sliding state; Indicates the absolute value of the response of the dynamic friction plate in the aircraft engine system; Indicates the contact stiffness of the static friction plate and the dynamic friction plate; Indicates the installation stiffness of the static friction plate; Indicates the mass of the static friction plate; represents the coefficient of friction; Indicates positive pressure.

[0072] like Figure 7As shown, the master-controlled elastic dry friction damper in this embodiment includes a dynamic friction plate 1, a sealing ring 2, a mounting ring 3, a static friction plate 4, a piezoelectric ceramic actuator 5, a damper mounting cylinder 6, an engine support frame 7, an engine elastic support 8, and an engine bearing 9. The master-controlled elastic dry friction damper is installed as follows: the dynamic friction plate 1 is connected to the engine elastic support 8 by bolts or other means to transmit vibration; the sealing ring 2 is used to prevent lubricating oil from entering the damper, thereby affecting the friction effect; the mounting ring has a three-ring spoke structure, which is used to mount the static friction plate 4 and also fix it to prevent twisting; the piezoelectric ceramic actuator 5 is a packaged product and is placed in the damper mounting cylinder 6 to provide the master control force; the damper mounting cylinder 6 has an annular groove structure, which is used to support and contain the pressure-protected piezoelectric ceramic actuator 5; the engine elastic support 8 is a folded-back integrated structure, with the engine bearing 9 mounted on one side to transmit bearing vibration and the other side connected to the engine support frame 7; the engine bearing 9 is mounted in the engine elastic support 8, which is then mounted to the engine support frame 7 via a flange. The piezoelectric ceramic actuator 5 is mounted in the damper mounting tube 6, which is mounted on the engine load-bearing frame 7. The static friction plate 4 is connected to the mounting ring 3, which is connected to the damper mounting tube 6. The sealing ring 2 is mounted in the mounting ring 3, and the dynamic friction plate 1 is bolted to one end of the engine elastic support 8.

[0073] Among them, based on the main control type elastic support dry friction damper model and the rotor maneuvering flight model, the aircraft engine rotor maneuvering flight dynamics model with the main control type elastic support dry friction damper is constructed as follows:

[0074] (4)

[0075] Where, is the rotor speed, is the angular velocity component of the rotor in the horizontal x direction, which is affected by the maneuvering flight action. is the system mass matrix; is the system damping matrix; is the system stiffness matrix; is the system gyro torque matrix; is the parameter stiffness matrix related to the rotor speed and the angular velocity component of the rotor angular velocity in the horizontal x direction; For external forces; is the rotor vibration amplitude; is the rotor vibration speed; is the rotor vibration acceleration; is the equivalent stiffness of the damper, Pick or , is the equivalent stiffness of the damper; is the equivalent damping of the damper, Pick or .

[0076] Step 22: The steps of obtaining the rotor vibration amplitude of the aircraft engine are as follows:

[0077] By solving Equation (4) using the numerical integration method, the rotor vibration amplitude of the air engine rotor under the control of the main control type elastic support dry friction damper can be obtained.

[0078] S3. Obtain the rotor vibration amplitude error at each moment;

[0079] Specifically, the amplitude extraction method is used to separate the single-frequency amplitude component, double-frequency amplitude component and triple-frequency amplitude component of the rotor vibration amplitude; based on the maximum amplitude component among the single-frequency amplitude component, double-frequency amplitude component and triple-frequency amplitude component and the reference rotor amplitude, the rotor vibration amplitude error at each moment is obtained.

[0080] Step 31, the steps of separating the single frequency amplitude component, the double frequency amplitude component and the triple frequency amplitude component of the rotor vibration amplitude are as follows:

[0081] In this embodiment, for the signal of the rotor vibration amplitude changing with time, the signal corresponding to the rotor vibration amplitude is input into the control system, and the Figure 5 The amplitude extraction method of the recursive least square algorithm (RLS) shown in the figure can separate the single frequency amplitude component, double frequency amplitude component and triple frequency amplitude component of the rotor vibration amplitude. That is, the time variation information of different double frequency amplitude components can be extracted, that is, the rotor vibration amplitude signal is first obtained. , let it be the input, and then set the filter parameters , initialize the RLS parameters and reference signals in sequence, and finally calculate and estimate the rotor vibration amplitude error vector.

[0082] Specifically, the amplitude extraction steps of the recursive least squares algorithm (RLS) are:

[0083] First, get the rotor vibration amplitude signal , let it be used as input; then set the filter parameters Initialize the RLS parameters and initialization parameters in sequence, and then calculate the estimated rotor vibration amplitude error vector. The error vector expression is:

[0084] (5)

[0085] in, is the expected amplitude sequence; is the amplitude estimation sequence; represents the estimated value; n is the time.

[0086] Then the calculation results are output. The result output is shown in formula (7):

[0087] (7)

[0088] in, is the input rotor vibration amplitude signal; is the filter coefficient vector, is the Kalman gain vector, and the superscript t is the matrix transpose symbol.

[0089] The error and correlation vectors or matrices are then updated, and the rotor amplitude can be extracted when the number of iterations is reached.

[0090] Step 32: Obtaining the rotor vibration amplitude error at each moment is as follows:

[0091] Multiply the single-frequency amplitude component, the double-frequency amplitude component and the triple-frequency amplitude component by the corresponding preset weights (the weights are designed as needed) to obtain the target single-frequency amplitude component, the target double-frequency amplitude component and the target triple-frequency amplitude component. Then, subtract the maximum amplitude component among the target single-frequency amplitude component, the target double-frequency amplitude component and the target triple-frequency amplitude component from the preset reference amplitude to obtain the rotor vibration amplitude error of the rotor at each moment.

[0092] S4. Obtaining the working conditions of the maneuvering flight;

[0093] Specifically, the overload coefficient of the aircraft during maneuvering flight is obtained according to the roll angle, and the operating conditions of the maneuvering flight are obtained according to the overload coefficient and a preset overload coefficient threshold, which include normal operating conditions, non-emergency operating conditions and emergency operating conditions.

[0094] Step 41: The overload factor of the aircraft during maneuvering flight is:

[0095] (8)

[0096] Where, is the roll angle, is the secant function; It is the overload factor of the aircraft during maneuvering flight.

[0097] Step 42: The step of obtaining the maneuvering flight condition based on the overload coefficient and the preset overload coefficient threshold range is as follows:

[0098] The preset overload coefficient thresholds include: a first overload coefficient threshold, a second overload coefficient threshold, and a third overload coefficient threshold, and the first overload coefficient threshold is smaller than the second overload coefficient threshold and smaller than the third overload coefficient threshold, such as Figure 6 The relationship between roll angle and overload factor is shown in the figure. During normal circling, the overload factor is only related to the roll angle, and the larger the roll angle, the greater the overload factor. When the roll angle exceeds 60°, the overload increases sharply. If the roll angle is 90°, the overload will tend to infinity and it is impossible to achieve. Generally, when the roll angle is 60°, the overload factor is 2, and then the overload factor increases significantly. In addition, the maximum overload factor that the pilot can withstand in 20 seconds to 30 seconds is 2. Therefore, in this embodiment, the first overload factor threshold is set to 1, the second overload factor threshold is set to 2, and the third overload factor threshold is set to 5; the maneuvering flight conditions are divided into three situations, as shown in Table 1.

[0099] Table 1

[0100]

[0101] Based on Table 1, when the overload coefficient is greater than or equal to the first overload coefficient threshold 1 and less than the second overload coefficient threshold 2, the aircraft engine is in a normal operating condition; when the overload coefficient is greater than or equal to the second overload coefficient threshold 2 and less than the third overload coefficient threshold 5, the aircraft engine is in a non-emergency operating condition; when the overload coefficient is greater than or equal to the third overload coefficient threshold 5, the aircraft engine is in an emergency operating condition.

[0102] S5. Obtaining a corresponding pressure control signal and controlling a damper of the aircraft engine according to the operating conditions of the maneuvering flight;

[0103] Specifically, when the operating condition of the maneuvering flight is a normal operating condition, the damper of the aircraft engine is controlled according to the set positive pressure control signal; when the operating condition of the maneuvering flight is a non-emergency operating condition, the pressure control signal is output according to the rotor vibration amplitude error and the PI control method, and the damper of the aircraft engine is controlled according to the pressure control signal; when the operating condition of the maneuvering flight is an emergency operating condition, the pressure control signal is output according to the rotor vibration amplitude error and the MFAC control method, and the damper of the aircraft engine is controlled according to the pressure control signal.

[0104] In this embodiment, when the working condition of the maneuvering flight is normal, the passive control method is selected. At this time, the control output is a positive pressure control signal, that is,

[0105] (9)

[0106] in, is the positive pressure control signal output at different times when the working condition of the maneuvering flight is the normal working condition, Indicates a constant pressure value.

[0107] In this embodiment, when the operating condition of the maneuvering flight is a non-emergency operating condition, the active control PI control method is adopted, that is:

[0108] (10)

[0109] Where, For non-emergency conditions n Pressure control signal at each moment; is the proportional coefficient of PI control; is the integral coefficient of PI control; For the n The rotor vibration amplitude error at each moment.

[0110] In this embodiment, when the operating condition of the maneuvering flight is an emergency condition, the MFAC control method of adaptive control is adopted, that is:

[0111] (11)

[0112] Where, For emergency situations n Pressure control signal at each moment; is the partial derivative; is the weight coefficient; is the step length; is the rotor vibration amplitude error between the ideal vibration amplitude and the actual vibration amplitude at the n+1th moment; is the number of iterations.

[0113] At this point, the main control spring-loaded friction damper changes the friction force according to the pressure control signal received to provide additional damping. and additional stiffness , thereby controlling the rotor vibration and reducing the vibration amplitude .

[0114] It should be noted that the active control method for aircraft engine rotor maneuvering flight according to the present invention can be designed Figure 2 The control system shown in FIG. 3 includes three parts: a tracking filter, a recursive least squares (RLS) module, and an adaptive control algorithm module. Collect amplitude signals for the rotor, 、 、 is the amplitude signal extracted by the filter, 、 、 is the specific amplitude obtained by RLS, 、 、 For different frequencies, To control the speed in the range, is the amplitude error, is the voltage. First, the rotor displacement signal collected Through the tracking filter, the speed frequency and frequency multiplication components in the signal are extracted in real time 、 、 Then the extracted signal is passed through the RLS module of the corresponding frequency to extract the amplitude of the corresponding frequency signal in real time. 、 、 Then, by comparing the amplitudes of each frequency, the maximum value is selected to solve the error with the reference amplitude. When the current speed is in the speed control range The error signal is Input adaptive control algorithm module to realize control voltage The output then drives the main control elastic dry friction damper to control the rotor vibration.

[0115] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the scope of protection of the present invention.

Claims

1. A method for active control of aircraft engine rotor maneuvering flight, characterized in that: include: Obtaining system attribute parameters of the aircraft engine and operating status parameters of the aircraft engine; The operating state parameters include: the common working line of the engine rotor, i.e., the rotation speed of each operating condition and the related maneuvering flight action type, wherein the maneuvering flight action type includes: turning and circling, diving and pulling up; Based on the system attribute parameters and operating state parameters of the aircraft engine, a rotor maneuvering flight dynamics model of an aircraft engine with a master-controlled elastic support friction damper is established. Based on the rotor maneuvering flight dynamics model of an aircraft engine with a master-controlled elastic support friction damper, the rotor vibration amplitude of the aircraft engine is obtained. The system attribute parameters of the aircraft engine include: system stiffness matrix, system damping matrix, system mass matrix, and system gyroscopic torque matrix. Establishing the rotor maneuvering flight dynamics model of an aircraft engine with a master-controlled elastic support friction damper includes: A rotor maneuvering flight model of an aircraft engine is constructed based on the rotor speed, the angular velocity component in the x-axis direction, the system attribute parameters of the aircraft engine, and the rotor vibration amplitude, rotor vibration velocity, and rotor vibration acceleration. The main control type elastic support dry friction damper model is adopted as the main control type elastic support dry friction damper dynamic model; Based on the main control type elastic support dry friction damper model and the rotor maneuvering flight model, the aircraft engine rotor maneuvering flight dynamics model with the main control type elastic support dry friction damper is constructed; The amplitude extraction method is used to separate the single-frequency amplitude component, double-frequency amplitude component, and triple-frequency amplitude component of the rotor vibration amplitude; the rotor vibration amplitude error at each moment is obtained based on the maximum amplitude component among the single-frequency amplitude component, double-frequency amplitude component, and triple-frequency amplitude component and the reference rotor amplitude; Obtain the overload factor of the aircraft during maneuvering flight based on the roll angle, and obtain the operating conditions of the maneuvering flight based on the overload factor and the preset overload factor threshold. The operating conditions include normal operating conditions, non-emergency operating conditions, and emergency operating conditions; When the operating condition of the maneuvering flight is a normal operating condition, the damper of the aircraft engine is controlled according to the set positive pressure control signal; when the operating condition of the maneuvering flight is a non-emergency operating condition, the pressure control signal is output according to the rotor vibration amplitude error and the PI control method, and the damper of the aircraft engine is controlled according to the pressure control signal; when the operating condition of the maneuvering flight is an emergency operating condition, the pressure control signal is output according to the rotor vibration amplitude error and the MFAC control method, and the damper of the aircraft engine is controlled according to the pressure control signal.

2. The method for active control of aircraft engine rotor maneuvering flight according to claim 1, characterized in that: The expression of the rotor maneuver flight model of an aircraft engine is: Where, is the rotor speed, is the angular velocity component of the rotor in the horizontal x direction, is the system mass matrix; is the system damping matrix; is the system stiffness matrix; is the system gyro torque matrix; is the parameter stiffness matrix related to the rotor speed and the angular velocity component of the rotor angular velocity in the horizontal x direction; For external forces; is the rotor vibration amplitude; is the rotor vibration speed; is the rotor vibration acceleration.

3. The active control method for aircraft engine rotor maneuvering flight according to claim 1, characterized in that: Obtaining the master-controlled elastic dry friction damper model includes: The expression of the main control type elastic support dry friction damper model in the viscous state is: The expression of the main control type elastic dry friction damper model in the slip state is: Where, represents the equivalent stiffness of the damper in the viscous state; It represents the equivalent stiffness of the damper in the slip state; represents the equivalent damping of the damper in the viscous state; It represents the equivalent damping of the damper in the sliding state; Indicates the absolute value of the response of the dynamic friction plate in the aircraft engine system; Indicates the contact stiffness of the static friction plate and the dynamic friction plate; Indicates the installation stiffness of the static friction plate; Indicates the mass of the static friction plate; represents the coefficient of friction; Indicates positive pressure.

4. The method for active control of aircraft engine rotor maneuvering flight according to claim 1, characterized in that: The expression of the flight dynamics model of an aero-engine rotor maneuver with a main-controlled elastic dry friction damper is: Where, is the rotor speed, is the angular velocity component of the rotor in the horizontal x direction, is the system mass matrix; is the system damping matrix; is the system stiffness matrix; is the system gyro torque matrix; is the parameter stiffness matrix related to the rotor speed and the angular velocity component of the rotor angular velocity in the horizontal x direction; For external forces; is the rotor vibration amplitude; is the rotor vibration speed; is the rotor vibration acceleration; is the equivalent stiffness of the damper, Pick or , is the equivalent stiffness of the damper; is the equivalent damping of the damper, Pick or .

5. The method for active control of aircraft engine rotor maneuvering flight according to claim 1, characterized in that: In non-emergency conditions, the expression of the pressure control signal is: Where, is the pressure control signal at the nth moment in non-emergency conditions; is the proportional coefficient of PI control; is the integral coefficient of PI control; is the rotor vibration amplitude error at the nth moment.

6. The method for active control of aircraft engine rotor maneuvering flight according to claim 1, characterized in that: In emergency conditions, the expression of the pressure control signal is: Where, is the pressure control signal at the nth moment in the emergency condition; is the partial derivative; is the weight coefficient; is the step length; is the rotor vibration amplitude error between the ideal vibration amplitude and the actual vibration amplitude at the n+1th moment; is the number of iterations.

7. The method for active control of aircraft engine rotor maneuvering flight according to claim 1, characterized in that: The steps for obtaining the maneuvering flight condition based on the overload factor and the preset overload factor threshold range are as follows: The preset overload coefficient thresholds include: a first overload coefficient threshold, a second overload coefficient threshold, and a third overload coefficient threshold, and the first overload coefficient threshold is smaller than the second overload coefficient threshold and smaller than the third overload coefficient threshold; When the overload factor is greater than or equal to the first overload factor threshold and less than the second overload factor threshold, the aircraft engine is in a normal operating condition; When the overload factor is greater than or equal to the second overload factor threshold and less than the third overload factor threshold, the aircraft engine is in a non-emergency operating condition; When the overload factor is greater than or equal to the third overload factor threshold, the aircraft engine is in an emergency operating condition.

8. The method for active control of aircraft engine rotor maneuvering flight according to claim 1, characterized in that: The RLS algorithm is used to separate the single frequency amplitude component, double frequency amplitude component and triple frequency amplitude component of the rotor vibration amplitude.

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