Active control method for maneuvering flight of aero-engine rotor

Abstract

The invention relates to the technical field of rotor vibration control, in particular to an aero-engine rotor maneuvering flight active control method, which comprises the following steps: acquiring system attribute parameters of an aero-engine and operation state parameters of the aero-engine; constructing an aero-engine rotor maneuvering flight dynamics model with a master control type elastic support dry friction damper, and obtaining a rotor vibration amplitude of the aero-engine; obtaining a rotor vibration amplitude error at each moment; acquiring the working condition of maneuvering flight; and according to the working condition of maneuvering flight, a corresponding pressure control signal is obtained, and a damper of the aero-engine is controlled. According to the invention, the vibration of the rotor is accurately controlled in a targeted and quick-response manner.

Description

Technical Field

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

[0002] Aircraft engines are a comprehensive reflection of a country's industrial foundation, scientific and technological level, and comprehensive national strength, and they play a very important role. In order to meet the requirements of high performance and high reliability of modern engines, the rotor system of the engine is gradually developing towards high speed and soft structure. Modern high-speed aircraft engines generally use rigid rotors mounted on flexible supports - dual-rotor engines with intermediate supports, with complex structures of intermediate bearings and multi-stage wheel blades (turbine disc blades, fan disc blades, compressor disc blades), which need to withstand high speed, dual-frequency excitation (high-pressure excitation and low-pressure excitation), strong coupling of high and low-pressure rotors, aerodynamic and thermal stresses, and special working conditions of maneuvering flight.

[0003] With the application and pre-research of the fourth-generation and fifth-generation aircraft, the maneuverability of the aircraft has become a key indicator in the evaluation system of new fighters. Typical maneuvers of aircraft include level flight acceleration, jump maneuvering flight, circling maneuvering flight, and somersault maneuvering flight. The high maneuverability of aircraft poses severe challenges to the design, operation and maintenance of aircraft engines. In a maneuvering flight environment, aircraft engines not only have high speed / acceleration, but also high maneuvering flight acceleration. Therefore, when the aircraft completes various maneuvering flight actions, in addition to being subjected to various excitations of its own, the engine rotor will also be subjected to the influence of additional excitation forces, that is, additional maneuvering inertia load excitation. These influences will cause the originally stable rotor system to suddenly produce large vibrations, which may in turn cause a series of faults such as rotor instability, friction between moving and static parts, and blade breakage.

[0004] In order to avoid the above phenomenon, the aircraft engine maneuvering flight needs to have a certain vibration control capability. The existing technology has some vibration control of vibration reduction and load reduction mechanisms, but most of them are passive control and it is difficult to fully meet the vibration control needs of all rotor mode maneuvering flights. 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] An active control method for aircraft engine rotor maneuvering flight of the present invention adopts the following technical solution, including: Obtaining system attribute parameters of the aircraft engine and operating state parameters of the aircraft engine; the operating state 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, wherein the maneuvering flight action type includes: turning and circling, diving and pulling up; According to the system attribute parameters and the operating state parameters of the aircraft engine, the aircraft engine rotor maneuvering flight dynamics model with a main control type elastic support dry friction damper is established; according to the aircraft engine rotor maneuvering flight dynamics model with a main control type elastic support dry friction damper, the rotor vibration amplitude of the aircraft engine is obtained; The single-frequency amplitude component, double-frequency amplitude component and triple-frequency amplitude component of the rotor vibration amplitude are separated by an amplitude extraction method; the rotor vibration amplitude error at each moment is obtained according to 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 overload factor of the aircraft during maneuvering flight is obtained according to the roll angle, and the working condition of the maneuvering flight is obtained according to the overload factor and the preset overload factor threshold, and the working condition includes normal working condition, non-emergency working condition and emergency working condition; When the operating condition of the maneuvering flight is a normal 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 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 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.

[0008] Preferably, the system property parameters of the aircraft engine include: a system stiffness matrix, a system damping matrix, a system mass matrix, and a system gyroscopic moment matrix.

[0009] Preferably, establishing a maneuvering flight dynamics model of an aeroengine rotor with a main control spring-supported friction damper comprises: A rotor maneuvering flight model of the aircraft engine is constructed 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, the rotor vibration velocity and the 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; 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.

[0010] Preferably, the rotor maneuvering flight model of the aircraft engine is expressed as:

[0011] In the formula, is the rotor speed, is the angular velocity component of the rotor angular velocity 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.

[0012] Preferably, obtaining the main control type elastic support dry friction damper model includes: The expression of the main control type elastic dry friction damper model in the viscous state is:

[0013] The expression of the main control type elastic dry friction damper model in the slip state is:

[0014] In the formula, 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; It represents 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 friction coefficient; Indicates positive pressure.

[0015] Preferably, the expression of the maneuvering flight dynamics model of an aeroengine rotor with a main control spring-supported dry friction damper is:

[0016] In the formula, is the rotor speed, is the angular velocity component of the rotor angular velocity 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 .

[0017] Preferably, in non-emergency conditions, the expression of the pressure control signal is:

[0018] In the formula, 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.

[0019] Preferably, in an emergency condition, the expression of the pressure control signal is:

[0020] In the formula, 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.

[0021] Preferably, the step of obtaining the working condition of the maneuvering flight according to the overload coefficient and the preset overload coefficient threshold range is: 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 condition.

[0022] 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.

[0023] The beneficial effects of the present invention are: 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 the main control elastic dry friction damper to the rotor maneuvering flight, and obtains the aircraft engine rotor maneuvering flight model with the main control elastic dry friction damper by combining it with the aircraft engine rotor maneuvering flight model. The aircraft engine rotor maneuvering flight model with the main control 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-multiplied 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, 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 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), H ∞, LQR (Linear Quadratic Control) method requires multiple adjustments and is difficult to meet the needs of rapid response. At the same time, 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. MFAC 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, and controls the damper of the aircraft engine according to the pressure control signal, so as to accurately control the rotor vibration in a targeted and fast-response manner. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. 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 creative work.

[0025] Figure 1 A flow chart of an active control method for aircraft engine rotor maneuvering flight according to the present invention; Figure 2 A control block diagram of an active control method for aircraft engine rotor maneuvering flight according to the present invention; Figure 3 It is a force analysis diagram of the turning and circling of the aircraft of the present invention; Figure 4 It is a force analysis diagram of the aircraft of the present invention when diving or pulling up; Figure 5 This is a flow chart of the amplitude extraction method based on the RLS algorithm; Figure 6 is a relationship diagram between the overload factor and the roll angle of the aircraft in the normal working condition of the present invention; Figure 7 It is a structural schematic diagram of the main control type elastic support dry friction damper of the present invention.

[0026] 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

[0027] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. 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 creative work are within the scope of protection of the present invention.

[0028] 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: S1. Obtaining system attribute parameters and operating status parameters of the aircraft engine; 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 types, among which the maneuvering flight action types include: turning and circling, diving and pulling up.

[0029] 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.

[0030] 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.

[0031] S2. Construct a maneuvering flight dynamics model of an aeroengine rotor with a main control spring-supported dry friction damper, and obtain the rotor vibration amplitude of the aeroengine; Specifically, a maneuvering flight dynamics model of an aircraft engine rotor with a main-controlled elastic-braced dry friction damper is established according to the system property parameters of the aircraft engine and the operating state parameters of the aircraft engine; and the rotor vibration amplitude of the aircraft engine is obtained according to the maneuvering flight dynamics model of an aircraft engine rotor with a main-controlled elastic-braced dry friction damper.

[0032] Step 21, the steps of establishing the rotor maneuvering flight dynamics model of an aircraft engine with a main-controlled elastic-bracket friction damper are as follows: construct 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 speed and rotor vibration acceleration; use the main-controlled elastic-bracket friction damper model as the main-controlled elastic-bracket friction damper dynamics model; construct the rotor maneuvering flight dynamics model of the aircraft engine with the main-controlled elastic-bracket friction damper according to the main-controlled elastic-bracket friction damper model and the rotor maneuvering flight model.

[0033] Among them, the differential equation of motion of the rotor system under different maneuvering flight states considered in this embodiment is: (1) In the formula, is the rotor speed, is the angular velocity component of the rotor angular velocity 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.

[0034] 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 recalculation of 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.

[0035] Among them, obtaining the main control type elastic support dry friction damper model includes: The expression of the main control type elastic dry friction damper model in the viscous state is: (2) The expression of the main control type elastic dry friction damper model in the slip state is: (3) In the formula, 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; It represents 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 friction coefficient; Indicates positive pressure.

[0036] like Figure 7 As shown, the main control type elastic support 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 bearing frame 7, an engine elastic support 8 and an engine bearing 9. Installation of the main control type elastic support dry friction damper: the dynamic friction plate 1 is connected to the engine elastic support 8 by bolts or other means for transmitting vibration; the sealing ring 2 is used to prevent the lubricating oil from entering the damper to affect the friction effect; the mounting ring is a 3-ring spoke structure for mounting the static friction plate 4 and can fix the static friction plate 4 for anti-twist; the piezoelectric ceramic actuator 5 is a packaged product, placed in the damper mounting cylinder 6, for providing the main control force; the damper mounting cylinder 6 is an annular groove structure for bearing the pressure-protecting electric ceramic actuator 5; the engine elastic support 8 is a return type integrated structure, one side of which is installed with the engine bearing 9 to transmit the vibration of the bearing, and the other side is connected with the engine bearing frame 7; the engine bearing 9 is installed in the engine elastic support 8, and the engine elastic support 8 is installed on the engine bearing frame 7 through a flange. The piezoelectric ceramic actuator 5 is installed in the damper mounting cylinder 6, the damper mounting cylinder 6 is installed on the engine load-bearing frame 7, the static friction plate 4 is connected to the mounting ring 3, and the mounting ring 3 is connected to the damper mounting cylinder 6. The sealing ring 2 is installed in the mounting ring 3, and the dynamic friction plate 1 is installed on one end of the engine elastic support 8 by bolts.

[0037] Among them, according to the main control type spring-supported dry friction damper model and the rotor maneuvering flight model, the aircraft engine rotor maneuvering flight dynamics model with the main control type spring-supported dry friction damper is constructed as follows: (4) In the formula, is the rotor speed, is the angular velocity component of the rotor angular velocity 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 .

[0038] Step 22: The steps of obtaining the rotor vibration amplitude of the aircraft engine are: 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 spring-supported dry friction damper can be obtained.

[0039] S3, obtaining the rotor vibration amplitude error at each moment; Specifically, the amplitude extraction method 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; based on the maximum amplitude component among the single-frequency amplitude component, the double-frequency amplitude component and the triple-frequency amplitude component and the reference rotor amplitude, the rotor vibration amplitude error at each moment is obtained.

[0040] 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: 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 Figure 5 The amplitude extraction method of the recursive least square algorithm (RLS) shown in FIG. 1 can separate the single-frequency amplitude component, double-frequency amplitude component and triple-frequency amplitude component of the rotor vibration amplitude. That is, the change information of different double-frequency amplitude components over time 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 turn, and finally calculate and estimate the rotor vibration amplitude error vector.

[0041] Specifically, the amplitude extraction steps of the recursive least squares algorithm (RLS) are: 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 turn, and then calculate and estimate the rotor vibration amplitude error vector. The expression of the error vector is: (5) in, is the expected sequence of amplitudes; is the amplitude estimation sequence; represents the estimated value; n is the time.

[0042] Then the calculation results are output. The result output is shown in formula (7): (7) 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.

[0043] The error and related vectors or matrices are then updated and the rotor amplitude can be extracted when the iteration number is reached.

[0044] Step 32, the step of obtaining the rotor vibration amplitude error at each moment is: Multiply the single-frequency amplitude component, double-frequency amplitude component and triple-frequency amplitude component by the corresponding preset weights (the weights are designed according to needs) to obtain the target single-frequency amplitude component, target double-frequency amplitude component and target triple-frequency amplitude component. Then, subtract the maximum amplitude component among the target single-frequency amplitude component, target double-frequency amplitude component and target triple-frequency amplitude component from the preset reference amplitude to obtain the rotor vibration amplitude error of the rotor at each moment.

[0045] S4. Obtaining the working conditions of the maneuvering flight; Specifically, the overload factor 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 factor and a preset overload factor threshold, and the operating conditions include normal conditions, non-emergency conditions and emergency conditions.

[0046] Step 41, the overload factor of the aircraft during maneuvering flight is: (8) In the formula, is the roll angle, is the secant function; It is the overload factor of the aircraft during maneuvering flight.

[0047] Step 42, the step of obtaining the working condition of the maneuvering flight according to the overload coefficient and the preset overload coefficient threshold range is: 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. Figure 6The relationship between the roll angle and the overload factor is shown in the figure. In normal circling, the overload factor is only related to the roll angle, and the larger the roll angle, the greater the overload factor. When it is greater than 60°, the overload increases sharply; if the roll angle is 90°, the overload will tend to infinity and it is impossible to achieve. In general, 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 s to 30 s 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.

[0048] Table 1

[0049] 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.

[0050] S5. According to the working condition of the maneuvering flight, obtain the corresponding pressure control signal and control the damper of the aircraft engine; Specifically, when the operating condition of the maneuvering flight is a normal 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 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 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.

[0051] In this embodiment, when the working condition of the maneuvering flight is normal, the passive control method is selected, and the control output is a positive pressure control signal, that is, (9) in, is the positive pressure control signal output at different times when the working condition of maneuvering flight is normal working condition, Indicates a constant pressure value.

[0052] In this embodiment, when the operating condition of the maneuvering flight is a non-emergency operating condition, the PI control method of active control is adopted, that is: (10) In the formula, 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 a certain moment.

[0053] 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: (11) In the formula, 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.

[0054] 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 .

[0055] 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 the figure includes three parts: tracking filter, recursive least squares (RLS) module and 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 is Through the tracking filter, the speed isochronous and multiplied frequency 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 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.

[0056] 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 principle of the present invention should be included in the protection scope of the present invention.

Claims

1. An active control method for 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 types, wherein the maneuvering flight action types include: turning and circling, diving and pulling up; According to the system attribute parameters and the operating state parameters of the aircraft engine, the aircraft engine rotor maneuvering flight dynamics model with a main control type elastic support dry friction damper is established; according to the aircraft engine rotor maneuvering flight dynamics model with a main control type elastic support dry friction damper, the rotor vibration amplitude of the aircraft engine is obtained; The single-frequency amplitude component, double-frequency amplitude component and triple-frequency amplitude component of the rotor vibration amplitude are separated by an amplitude extraction method; the rotor vibration amplitude error at each moment is obtained according to 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 overload factor of the aircraft during maneuvering flight is obtained according to the roll angle, and the working condition of the maneuvering flight is obtained according to the overload factor and the preset overload factor threshold, and the working condition includes normal working condition, non-emergency working condition and emergency working condition; When the operating condition of the maneuvering flight is a normal 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 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 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 system attribute parameters of aircraft engines include: system stiffness matrix, system damping matrix, system mass matrix, and system gyroscopic torque matrix.

3. The active control method for aircraft engine rotor maneuvering flight according to claim 2, characterized in that: The establishment of the flight dynamics model of an aeroengine rotor maneuver with a main control spring-loaded friction damper includes: A rotor maneuvering flight model of the aircraft engine is constructed 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, the rotor vibration velocity and the 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; 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.

4. The active control method for aircraft engine rotor maneuvering flight according to claim 3, characterized in that: The expression of the rotor maneuvering flight model of an aircraft engine is: In the formula, is the rotor speed, is the angular velocity component of the rotor angular velocity 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.

5. The method for active control of aircraft engine rotor maneuvering flight according to claim 3, characterized in that: Obtaining the master-controlled elastic dry friction damper model includes: The expression of the main control type elastic 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: In the formula, 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; It represents 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 friction coefficient; Indicates positive pressure.

6. The method for active control of aircraft engine rotor maneuvering flight according to claim 3, characterized in that: The expression of the maneuvering flight dynamics model of an aeroengine rotor with a main control spring-loaded friction damper is: In the formula, is the rotor speed, is the angular velocity component of the rotor angular velocity 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 .

7. The active control method for aircraft engine rotor maneuvering flight according to claim 1, characterized in that: In non-emergency conditions, the expression of the pressure control signal is: In the formula, 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.

8. 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: In the formula, 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.

9. The active control method for aircraft engine rotor maneuvering flight according to claim 1, characterized in that: The steps for obtaining the working condition of maneuvering flight according to 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 condition.

10. The active control method for 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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