A method, device and system for servo motor flutter stabilization control

Abstract

This application discloses a servo motor flutter stabilization control method, device, and system, relating to the field of aircraft servo motor control technology. The method includes: determining that the aircraft has entered the impact unloading stage and triggering an output disturbance request signal when the aircraft's acceleration signal meets the impact unloading condition; responding to the disturbance request signal, calculating the disturbance frequency based on a mechanical model, and obtaining the disturbance amplitude based on the maximum acceleration value by retrieving a matching gradation lookup table; synthesizing a disturbance sinusoidal signal based on the disturbance frequency and the disturbance amplitude; and linearly superimposing the disturbance sinusoidal signal with the servo motor current control signal to obtain a total current control signal; setting the disturbance sinusoidal signal to zero and recording a fault flag when the absolute value of the total current control signal is greater than a current threshold. By superimposing a disturbance sinusoidal signal on the servo motor current control signal, the control dead zone introduced by mechanical backlash can be effectively suppressed, ensuring the accuracy of the control surface following and the stability and timeliness of the aircraft's attitude control.

Description

Technical Field

[0001] This application relates to the field of aircraft servo motor control technology, and in particular to a servo motor flutter stabilization control method, device and system. Background Technology

[0002] In the control system of high-maneuverability aircraft, the servo mechanism is the core actuator for flight attitude control, and its reliability directly determines the flight performance and control accuracy of the aircraft. With the development of aerospace technology, during launch separation and high-G maneuvers, the servo mechanism needs to withstand extremely high instantaneous impact loads. The servo motor and the mounting interface will generate violent dynamic responses, which will lead to a decrease in the transmission accuracy of the servo motor over a long period of time, deteriorate the control performance, and cause safety hazards.

[0003] Currently, engineering solutions for ensuring the reliability of installation interface connections mainly fall into two categories. One approach involves increasing bolt size, raising preload, or using high-strength impact-resistant materials. This significantly increases structural weight, hindering the lightweight and compact installation requirements of aircraft. Furthermore, passive reinforcement cannot withstand instantaneous impact loads, and connection failure remains a risk under extreme conditions. The other approach involves introducing active devices such as piezoelectric ceramic actuators to compensate for micro-displacements. This requires additional installation space and complex drive circuits, increasing system complexity. Moreover, these devices have weak survivability under high-impact environments, leading to increased weight and cost.

[0004] However, under high impact loads, stress wave reflection can cause a sudden drop in bolt preload, leading to problems such as loosening of the force transmission chain, a surge in mechanical backlash, and deterioration of control accuracy. Existing servo control algorithms are mostly based on position loop or velocity loop feedback, which are difficult to respond within the millisecond or even microsecond window of high impact. Summary of the Invention

[0005] The main purpose of this application is to provide a servo motor flutter stabilization control method, device and system, which aims to solve the problem that when an aircraft is subjected to high overload impact environment, the preload of the servo motor mounting interface drops instantaneously due to stress wave reflection, which leads to the loosening of the force transmission chain, increased mechanical clearance and excessive control accuracy of the control surface. At the same time, it solves the problems of significant weight increase, slow response speed and reliance on additional actuation hardware in traditional passive reinforcement.

[0006] To achieve the above objectives, this application provides a servo motor flutter stabilization control method, applied to the servo motor controller of an aircraft, comprising: When the acceleration signal of the aircraft meets the impact unloading condition, it is determined that the aircraft has entered the impact unloading stage and triggers the output disturbance request signal; the acceleration signal is the acceleration value output by the acceleration sensor acquired at a fixed sampling frequency. In response to the disturbance request signal, the disturbance frequency is calculated based on the mechanical model, and the disturbance amplitude is obtained by retrieving a matching range lookup table based on the maximum acceleration value; the range lookup table is a correspondence between the intervals of acceleration values ​​and the disturbance amplitudes created based on the mechanical equilibrium equations. Based on the disturbance frequency and the disturbance amplitude, a disturbance sinusoidal signal is synthesized; and the disturbance sinusoidal signal is linearly superimposed with the servo motor current control signal to obtain the total current control signal. When the absolute value of the total current control signal is greater than the current threshold, the disturbance sinusoidal signal is set to zero and a fault flag is recorded; the current threshold is a preset multiple of the motor's rated current.

[0007] Optionally, when the acceleration signal of the aircraft meets the impact unloading condition, it is determined that the aircraft has entered the impact unloading phase, including: If the absolute value of the acceleration is greater than the acceleration threshold and the rate of change of acceleration is negative, the aircraft is determined to have entered the impact unloading phase. The acceleration threshold is dynamically determined based on the bolt preload and the equivalent mass of the transmission chain at the current installation interface.

[0008] Optionally, the calculation of the disturbance frequency based on the mechanical model includes: The disturbance frequency is calculated based on the following formula;

[0009] in, For the perturbation frequency, For the contact stiffness of the installation interface, The equivalent mass of the transmission chain. For safety reasons, This is the frequency offset.

[0010] Optionally, the method further includes: The disturbance amplitude A and the maximum acceleration are obtained based on the inverse formula of the mechanical equilibrium equation. The correspondence;

[0011] in, ξ For safety margin, m load This refers to the total load mass of the control surfaces and transmission mechanism. To achieve the maximum acceleration during the impact unloading phase, The equivalent mass of the transmission chain. For the perturbation frequency, The angular frequency of the disturbance signal; A grading comparison table is established based on the correspondence between the disturbance amplitude and the range of the maximum acceleration.

[0012] Optionally, after setting the disturbance sinusoidal signal to zero when the absolute value of the total current control signal is greater than the current threshold, the method further includes: When the event of setting the disturbance sinusoidal signal to zero occurs continuously for a preset number of times, the superposition of the disturbance sinusoidal signal on the servo current control signal is terminated.

[0013] Furthermore, to achieve the above objectives, this application also provides a servo flutter stabilization control device, integrated within the servo controller of an aircraft, the device comprising: The impact detection module determines that the aircraft has entered the impact unloading phase and triggers the output disturbance request signal when the acceleration signal of the aircraft meets the impact unloading conditions; the acceleration signal is the acceleration value output by the acceleration sensor acquired at a fixed sampling frequency. The parameter calculation module, in response to the disturbance request signal, calculates the disturbance frequency based on the mechanical model, and obtains the disturbance amplitude based on the maximum acceleration by retrieving a matching range lookup table; the range lookup table is a correspondence between the intervals of acceleration values ​​and the disturbance amplitudes created based on the mechanical equilibrium equations. The signal synthesis module synthesizes a disturbance sinusoidal signal based on the disturbance frequency and the disturbance amplitude; and linearly superimposes the disturbance sinusoidal signal with the servo motor current control signal to obtain the total current control signal. The protection monitoring module sets the disturbance sinusoidal signal to zero and records the fault flag when the absolute value of the total current control signal is greater than the current threshold; the current threshold is a preset multiple of the motor's rated current.

[0014] Optionally, the protection monitoring module is further configured to terminate the superposition of the disturbance sinusoidal signal on the servo motor current control signal when the event of the disturbance sinusoidal signal being set to zero is detected to reach a preset number of consecutive times.

[0015] In addition, this application also provides a servo motor flutter stabilization control system, including: a single-axis acceleration sensor, a servo motor flutter stabilization control device as described in the foregoing embodiments, a current loop, a power drive unit, and a servo motor. The single-axis accelerometer is connected to the input of the servo flutter stabilization control device to provide it with the aircraft's acceleration signal; the output of the servo flutter stabilization control device is connected to the input of the current loop; the output of the current loop is connected to the power drive unit; and the power drive unit is used to drive the servo motor.

[0016] This application proposes a servo motor flutter stabilization control method, device, and system. By superimposing a micro-disturbance sinusoidal signal synchronized with the impact state onto the current loop of the servo motor controller, and when the impact unloading condition is met, the disturbance signal is calculated based on a mechanical model and superimposed onto the servo motor current control signal. The inertial force generated by the equivalent mass of the transmission chain dynamically compensates for the interface pressure loss during the impact unloading stage. This solves the problems of instantaneous drop in preload force at the servo motor mounting interface, loosening of the transmission chain, and deterioration of control accuracy under high overload impact. It achieves millisecond-level reliable contact of the interface transmission chain without additional hardware, effectively suppresses the control dead zone introduced by mechanical backlash, and ensures the servo surface following accuracy and the stability of aircraft attitude control. Attached Figure Description

[0017] Figure 1 This is a flowchart illustrating a servo motor flutter stabilization control method according to an embodiment of this application. Figure 2 This is a flowchart illustrating a specific method for controlling servo motor flutter stabilization according to an embodiment of this application. Figure 3 This is a timing diagram illustrating the interface pressure compensation principle of a servo motor flutter stabilization control method according to an embodiment of this application. Figure 4 This is a schematic diagram of the architecture and signal flow of a servo motor flutter stabilization control system according to an embodiment of this application; Figure 5 This is a structural block diagram of a servo motor flutter stabilization control device according to an embodiment of this application.

[0018] The realization of the purpose, functional features and advantages of this application will be further explained in conjunction with the embodiments and with reference to the accompanying drawings. Detailed Implementation

[0019] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative effort are within the scope of protection of the present invention.

[0020] In related technologies, in the control system of high-maneuverability aircraft, under high overload impact conditions, stress waves are reflected at the interface between the servo and the aircraft structure, which causes the preload of the mounting bolts to drop instantaneously during the impact unloading phase, thereby causing the force transmission chain to loosen, the servo body and the support to separate microscopically, and interrupting the interface friction force transmission; at the same time, nonlinear mechanical backlash is introduced, causing dead zones in the control surface, directly causing control surface following errors, and deteriorating control stability and accuracy.

[0021] Existing technologies mostly employ passive reinforcement or additional actuation devices for compensation. The former significantly increases weight, while the latter has a slow response and relies on additional hardware, making it difficult to meet engineering requirements. Therefore, this invention proposes a dynamic force transmission chain maintenance control method for the servo motor mounting interface under high overload impact. Based on the principle of high-frequency inertial force dynamic compensation, a high-frequency micro-disturbance signal is superimposed on the current loop feedforward. The inertial force generated by the equivalent mass of the transmission chain offsets the impact unloading force, achieving millisecond-level reliable contact maintenance of the interface force transmission chain without additional hardware, ensuring control accuracy. Based on the above working principle, the technical effect of this application's solution is physically certain and inevitable: the essence of interface pressure loss during the impact unloading stage is that the direction of external load acceleration is opposite to the direction of bolt preload, resulting in an instantaneous decrease in effective contact pressure; while the high-frequency micro-disturbance applied in this solution, driven by a servo motor, directly generates a dynamic inertial force with the opposite direction and matching magnitude to the pressure loss using the equivalent mass of the transmission chain itself. The response speed of this compensation force is only limited by the current loop bandwidth and the natural frequency of the mechanical system.

[0022] Figure 1 This is a flowchart illustrating a servo motor flutter stabilization control method according to an embodiment of this application.

[0023] like Figure 1 The diagram illustrates a servo motor flutter stabilization control method according to an embodiment of this application; applied to an aircraft servo motor controller; the method includes: Step S101: When the acceleration signal of the aircraft meets the impact unloading condition, it is determined that the aircraft has entered the impact unloading stage and triggers the output disturbance request signal; the acceleration signal is the acceleration value output by the acceleration sensor collected at a fixed sampling frequency.

[0024] It should be noted that the aircraft described is a high-maneuverability, high-G aircraft, such as tactical missiles and guided weapons. When an aircraft transitions from high-G maneuvers (such as dives and climbs) to level flight, the servo mechanical structure experiences enormous alternating stress due to sudden load changes, falling under the condition of interface structural fatigue damage. The impact unloading phase occurs after the aircraft has undergone a high-G impact, reaching its peak acceleration and beginning to decline, during which structural stress waves are reflected. The servo controller is a high-performance electric servo controller used to suppress active stress. The acceleration signal is the signal value output from the aircraft's single-axis acceleration sensor.

[0025] In the embodiments of this application, the acceleration signal of the aircraft is acquired in real time by an accelerometer. The system determines whether the acceleration signal meets the impact unloading conditions. If the conditions are met, the aircraft enters the impact unloading phase and triggers an output disturbance request signal. The impact unloading conditions include an acceleration amplitude exceeding an acceleration threshold and a negative rate of change of acceleration. If the aircraft control system detects that the impact unloading conditions are met, it needs to actively intervene in the dangerous situation and trigger an output disturbance request signal. This ensures that the disturbance signal is activated only when most needed, avoiding unnecessary energy consumption and interference with normal control.

[0026] Step S102: In response to the disturbance request signal, calculate the disturbance frequency based on the mechanical model, and obtain the disturbance amplitude based on the maximum acceleration by retrieving a matching graded lookup table; the graded lookup table is the correspondence between the interval of acceleration values ​​and the disturbance amplitude created based on the mechanical equilibrium equation.

[0027] It should be noted that the maximum acceleration value is the peak acceleration collected during the impact unloading phase, corresponding to the magnitude of the impact overload. The mechanical model is a frequency calculation model established based on the contact stiffness of the installation interface, the safety factor, and the frequency deviation. The grading table is a pre-established mapping table of different ranges of impact acceleration values ​​and their corresponding disturbance amplitudes, based on the mechanical equilibrium equations. The mechanical equilibrium equations describe the balance relationship between pressure loss and disturbance compensation inertial force during the impact unloading phase.

[0028] In embodiments of this application, in response to the disturbance request signal, the disturbance frequency can be calculated based on equation (1) of a mechanical model (typically related to the natural modal frequencies and equivalent mass of the servo / control surface). (1) in, The frequency of the disturbance at the contact interface. For the contact stiffness of the installation interface, The equivalent mass of the transmission chain. For safety reasons, This refers to the frequency offset. A disturbance frequency that matches the stress release frequency can be selected. By increasing the frequency offset, it avoids the fundamental frequency of the aircraft's servo mounting structure without exceeding the servo motor bandwidth. The disturbance amplitude is obtained by retrieving a pre-stored range table. This range table is pre-calculated based on the mechanical equilibrium equations, mapping different acceleration amplitude ranges to corresponding disturbance amplitudes. This ensures the targeting of the disturbance signal, precise frequency matching of the system's dynamic characteristics, and adaptive adjustment of the amplitude according to the impact intensity, thereby achieving stress suppression and avoiding overcompensation or undercompensation.

[0029] Step S103: Based on the disturbance frequency and the disturbance amplitude, a disturbance sinusoidal signal is synthesized; and the disturbance sinusoidal signal is linearly superimposed with the servo motor current control signal to obtain the total current control signal.

[0030] It should be noted that the servo current control signal originates from the reference control signal of the descent flight master control unit of the aircraft control system. Calculated by the flight controller based on attitude sensors (gyroscopes, accelerometers, attitude calculations), flight mission commands, and attitude control laws, it is sent to the servo controller and serves as the servo reference target current control command.

[0031] In embodiments of this application, the aforementioned perturbation frequency can be used as a basis. The disturbance amplitude A is processed by DDS (Direct Digital Synthesizer) to generate a standard sinusoidal disturbance signal. The high-frequency micro-perturbation signal is a small-amplitude, specific-frequency alternating signal; the perturbation sinusoidal signal is linearly superimposed with the servo motor current control signal by a built-in adder to obtain the total current control signal. ,in, This is the normal servo current control signal, which is also the servo reference target position signal. Therefore, it can be seen that linearly superimposing the servo reference target position signal with a high-frequency micro-disturbance sinusoidal signal generates the total current control signal, maintaining reliable contact of the force transmission chain at the mounting interface while ensuring normal attitude control of the aircraft.

[0032] Step S104: When the absolute value of the total current control signal is greater than the current threshold, the disturbance sinusoidal signal is set to zero and a fault flag is recorded; the current threshold is a preset multiple of the motor's rated current.

[0033] It should be noted that the rated current of the motor is the maximum safe operating current that the servo motor is allowed to flow continuously under standard operating conditions. When the operating current does not exceed the rated current, the motor temperature rise is normal, and there is no overheating or coil burnout; when the operating current exceeds the rated current, overheating will increase, magnets will demagnetize, insulation will age, and the motor and drive circuit may even burn out.

[0034] In the embodiments of this application, a limiting check operation can be performed. When the absolute value of the total current control signal of the servo motor is detected to be greater than a current threshold, such as a preset multiple of the motor's rated current, the disturbance sinusoidal signal is set to zero, retaining only the normal servo motor current control signal, and a fault identifier is recorded. If the event of setting the disturbance sinusoidal signal to zero occurs consecutively for a preset number of times, such as three times, the superposition of the disturbance sinusoidal signal on the servo motor current control signal is terminated. This prevents excessive current after the disturbance sinusoidal signal is superimposed, returning to the safe basic control mode and protecting the servo motor. The current threshold is a preset multiple of the servo motor's rated current, such as 1.2 times or 1.5 times. Therefore, it can be ensured that while a disturbance signal is injected, the total current is prevented from exceeding the safe operating range of the servo motor, avoiding equipment damage.

[0035] For example, a high-range single-axis accelerometer connected to the aircraft control system can be used to collect high-frequency acceleration signals of the aircraft in real time during launch separation and high-G maneuvers. The acquisition frequency is set to 1MHz to ensure that instantaneous signal changes during the impact process are captured; the rate of change of the acquired high-frequency acceleration signal can be calculated in real time by the processing unit inside the servo controller. Impact unloading conditions are and At that time, it was determined that the aircraft's servo motors had entered the impact unloading phase. Among these, The acceleration threshold can be determined based on the bolt preload at the current installation interface. Equivalent mass of transmission chain Dynamically determined. For example, the formula. ,in, The rated capacity is 5000N, and the equivalent mass of the transmission chain is... The calibrated value is 0.2 kg, and the acceleration threshold is calculated. It is 12500g. For example, in When the g = 15000g, the aircraft is determined to have entered the impact unloading phase, triggering the output disturbance request signal.

[0036] Then, based on the locked impact state, the disturbance frequency can be calculated using the following mechanical model (1). ,in, For the contact stiffness of the installation interface, For safety reasons, Frequency offset. Contact stiffness of the mounting interface. The calibration value is 3.125 × 10¹² N / m, with a safety factor of 1.5%. The value is 0.006, which represents the frequency offset. The value is taken as 50Hz, and the frequency is calculated. Locked at 800 Hz, this frequency avoids both the fundamental frequency of the servo mounting structure (500 Hz) and the bandwidth of the servo motor (1200 Hz), ensuring sufficient compensating inertial force without triggering structural resonance. The maximum acceleration value can be identified using a comparator, and the disturbance amplitude A can be obtained by retrieving the matching range lookup table based on this maximum acceleration value. The range lookup table is a table of correspondence between acceleration values ​​and disturbance amplitudes calculated in advance based on the mechanical equilibrium equation (2): (2) in, F inertial For inertial compensation force, To account for the maximum pressure loss during the impact unloading phase, among which, The stress wave attenuation coefficient can take values ​​of [0.3, 0.5]. For transmission efficiency. The disturbance amplitude A can be obtained by inverse solution of the variant form (3): (3) in, ξ For safety margin, m load The total load mass of the control surface and transmission mechanism; This represents the maximum acceleration during the impact unloading phase. The equivalent mass of the transmission chain; The perturbation frequency; This is the angular frequency of the disturbance signal. For example, an impact overload of 15000 g corresponds to an amplitude of 3. μ m.

[0037] In servo motor flutter control methods, the calculated high-frequency micro-perturbation sinusoidal signal can be converted using a digital adder. With normal servo current control signal Linear superposition is performed to generate the total current control signal. It drives the servo motor through the subsequent current loop and power drive unit.

[0038] Simultaneously, it can detect the current value corresponding to the total current control signal of the servo motor in real time, and the absolute value of the total current control signal... At that time, Set to zero and record the fault flag; where, This is the motor's rated current. If an overcurrent is detected within three consecutive PWM cycles, a lockout signal is immediately output. This stops the superposition of the high-frequency micro-disturbance signal onto the normal servo current control signal and sets the high-frequency micro-disturbance signal to zero, retaining only the normal flight control commands to prevent the motor from overheating or being damaged.

[0039] Therefore, the current increment introduced by the superimposed high-frequency micro-disturbance signal in this embodiment is approximately 5% of the motor's rated current, far below the 110% safety threshold. Thus, the disturbance signal can be continuously injected for about 50 ms, covering the entire impact unloading process. This solves the problems of instantaneous drop in preload force at the servo mounting interface, loosening of the transmission chain, and deterioration of control accuracy under high overload impacts. It achieves millisecond-level reliable contact of the interface transmission chain without additional hardware, effectively suppressing the control dead zone introduced by mechanical backlash and ensuring the servo surface following accuracy and aircraft attitude control stability.

[0040] Figure 2 This is a flowchart illustrating a specific method for controlling servo motor flutter stabilization according to an embodiment of this application.

[0041] like Figure 2 The diagram illustrates the specific flow of a servo motor flutter stabilization control method, applied to the dynamic force chain holding control scenario of the servo motor mounting interface in high-maneuverability aircraft. The flow of this servo motor flutter stabilization control method is shown below, with each step and process described in detail. Figure 1 One-to-one correspondence enables precise intervention during the impact unloading phase.

[0042] It should be noted that the grading table is based on the mechanical equilibrium equations and is a pre-established mapping table that maps different ranges of impact acceleration values ​​to their corresponding disturbance amplitudes. The mechanical equilibrium equations describe the balance relationship between pressure loss and disturbance compensation inertial force during the impact unloading phase.

[0043] After initiating the control process, the servo motor flutter stabilization control system enters real-time acceleration signal detection mode. During this phase, a high-frequency single-axis accelerometer was used at a sampling rate of 500 kHz to continuously acquire acceleration sensor signals from the aircraft during launch or high-G maneuvers. The acceleration signals were then analyzed. After performing a moving average filter (e.g., with a window length of 5 sampling points).

[0044] Perform impact unloading judgment; first, perform judgment 1: | Is it greater than? Among them, acceleration threshold It is based on the bolt preload at the installation interface. F pre =3125N, equivalent mass of the transmission chain =0.02kg, gravitational acceleration g≈10m / s² 2 It can be done Calculations show that =12500g; in Then, the rate of change of acceleration can be calculated. Otherwise, return to the previous level. If condition 1 is satisfied, proceed to condition 2: At that time, a disturbance trigger signal is output to "lock the unloading phase" in order to latch the current moment. To trigger the zero point, a hardware interrupt request is sent to the subsequent computing modules.

[0045] Upon receiving a disturbance trigger signal, the servo motor flutter stabilization control system enters the disturbance parameter calculation phase, which can calculate the disturbance frequency based on a mechanical model. and disturbance amplitude Equation (4) of the physical model for inertial force compensation can be used: (4) Obtain the perturbation frequency With contact stiffness k contact and equivalent quality m eg The quantitative relationship, in which ，k contact For the contact stiffness of the installation interface, The equivalent mass of the transmission chain. k For safety factors; and this can be achieved by increasing the frequency offset. Avoid the structural fundamental frequency and do not exceed the motor bandwidth.

[0046] The disturbance amplitude A and the peak impact value can also be obtained through the modified form (3). Proportional to, with Inversely proportional, that is, For example, using a grading table, impact overloads are divided into three levels: 5000g, 10000g, and 15000g, corresponding to amplitudes of 1... μ m、2 μ m and 3 μ m. The corresponding amplitude A can be directly retrieved, and then a micro-perturbation sinusoidal signal can be generated based on the perturbation frequency and perturbation amplitude. That is, the frequency is A sinusoidal signal with an amplitude of A. During the disturbance signal superposition stage, the generated micro-disturbance signal, such as a sinusoidal signal, can be superimposed on the main control signal. For example, the micro-disturbance sinusoidal signal and the normal servo current control signal from the main CPU of the flight control system can be linearly superimposed by the digital adder in the signal synthesis module to form the total current control signal, which is then output to the current loop.

[0047] Safety limiting logic can be embedded in the servo controller firmware to set the current threshold I output to the servo. maxSet to 5A. If the absolute value of the superimposed total current control signal exceeds this current threshold, the micro-disturbance signal will be forcibly set to zero, and a fault flag will be recorded; otherwise, the total current control signal will be directly output to control the power drive unit to output AC power, driving the servo motor to run, thereby completing all control processes.

[0048] Therefore, through the above process, the embodiments of this application only begin to inject high-frequency micro-disturbance signals when the aircraft enters the critical window of impact unloading. The inertial force generated by the equivalent mass of the transmission chain is used to offset the interface pressure loss. Reliable contact of the force transmission chain at the installation interface can be maintained without additional hardware, effectively solving the problem of mechanical clearance and deterioration of control accuracy caused by preload drop.

[0049] Figure 3 This is a timing diagram illustrating the interface pressure compensation principle of another servo motor flutter stabilization control method according to an embodiment of this application.

[0050] like Figure 3 As shown, a timing diagram of the interface pressure compensation principle of a servo motor flutter stabilization control method is provided. It can accurately offset the interface pressure loss during the impact unloading stage through the dynamic compensation principle of high-frequency inertial force.

[0051] It should be noted that the horizontal axis represents time ( t The vertical axis represents force. F or acceleration The upper coordinate system of the figure shows the curves of time versus acceleration; the lower coordinate system shows three curves. The long dashed line represents the impact overload curve; at time t0, the acceleration reaches its maximum value. ， It then rapidly falls back to the critical point of 0. The short dashed line segment represents the interface preload curve; the servo motor mounting interface preload curve, as acceleration increases, the peak value of the impact overload decreases, and during the impact unloading phase after t0, the interface preload drops rapidly, resulting in preload loss. At this point, the force transmission chain between the servo motor and the mounting bracket faces the risk of loosening. The solid curve represents the inertial compensation force curve for compensating for preload loss. The compensation force generated by the high-frequency micro-disturbance signal is injected only during the "golden window" when the preload drops (i.e., the stage at the beginning of impact unloading) to avoid false triggering under non-impact conditions.

[0052] Step 1: Precisely lock the impact unloading phase In the embodiments of this application, such as Figure 3 The triggering begins at the "golden window" indicated, precisely locking the impact unloading phase. Single-axis acceleration signals can be acquired in real-time using a high-range accelerometer on the aircraft at a fixed sampling frequency of 500Hz. This allows for the acquisition of impact overload curve data. Precise phase locking can be achieved by executing a dual-criteria judgment. In the first judgment, | can be detected. | Is it greater than the dynamic acceleration threshold? In this embodiment Based on the bolt preload at the installation interface F0=3125N, and the equivalent mass of the transmission chain... =0.02 kg, can be expressed by the following formula =0.8F0 / 0.02g, calculated as follows =12500g. Therefore, when | When |> 12500g, the second stage of judgment is initiated; during the second stage of judgment, the rate of change of the acceleration signal is calculated. ,when When the acceleration value has reached its maximum value and begins to fall back into negative territory, the aircraft's servos begin to enter the impact unloading phase. At this time, the aircraft's control system outputs a disturbance trigger signal to lock the impact unloading phase, precisely locking the compensation timing (i.e., the attached...). Figure 3 The "golden window" in the process is used to avoid false triggering under non-impact conditions.

[0053] Step 2: Calculate the disturbance frequency based on the mechanical model And the disturbance amplitude A.

[0054] In embodiments of this application, in response to the disturbance trigger signal, the high-frequency micro-disturbance frequency can be calculated based on a mechanical model. The disturbance amplitude A is as follows: At the disturbance frequency In the calculation process, to avoid structural resonance and adapt to the motor bandwidth, the mechanical model (1) can be used to calculate the disturbance frequency. As an example, For the contact stiffness of the installation interface, a safety factor is used. Frequency offset Equivalent mass of the transmission chain Calculated This frequency avoids the base frequency of the servo motor mounting structure. ( ),and ( The frequency is within a safe operating frequency band and does not exceed the servo motor bandwidth (e.g., 6 kHz), thus ensuring that the compensation force is effective and does not damage the structure. Based on the above model calculations, the disturbance frequency is approximately 900 Hz. Depending on the actual application, the disturbance frequency can be locked within a safe frequency band between 600 Hz and 1000 Hz.

[0055] In the calculation of the disturbance amplitude A, the required micro-displacement amplitude can be calculated based on the principle of inertial force compensation balance. That is, the disturbance amplitude A can be obtained by inverse solution based on the variant form (3).

[0056] In the embodiments of this application, the maximum acceleration value detected during the impact overload phase can be combined. The corresponding disturbance amplitude A is obtained by inverse kinematics transformation. The maximum acceleration of the impact overload is divided into three levels: (0~5000g), [5000g~10000g], and (10000g~15000g], which correspond to preset disturbance amplitudes of 1. μ m、2 μ m and 3 μ m. Based on the above correspondence between disturbance amplitude and maximum acceleration range, a tiered reference table is pre-established, which can be used based on real-time data acquisition. Determine the corresponding overload setting and directly retrieve the corresponding disturbance amplitude A. For example, in When the weight is 12000g, A=3. μ m. It can achieve a direct mapping from impact overload to disturbance amplitude without complex iterative calculations, and can be completed in a single cycle in an FPGA, meeting millisecond-level response requirements. Therefore, it can be achieved through the disturbance frequency. In conjunction with the amplitude A, it ensures sufficient compensation for inertial force while avoiding motor saturation due to excessive amplitude. Furthermore, by combining it with a preset grading lookup table, the calculation results can be stored in the grading lookup table in advance, enabling fast retrieval of LUT amplitude lookup table (LUT) method, thereby avoiding complex real-time mathematical calculations within the chip and significantly improving real-time performance.

[0057] Step 3: Generation and superposition of micro-perturbation signals It should be noted that the DDS module generates high-precision, programmable micro-disturbance signals for servo control of servo motors, enabling precise phase locking, active vibration reduction / elimination, and nonlinear compensation.

[0058] In embodiments of this application, a perturbation signal can be generated in the following manner. First, the perturbation parameters calculated in step 2 are obtained, for example... =900Hz and A=2 μ m, and then a high-frequency micro-perturbation signal is generated through the DDS module. At time t0, the impact acceleration reaches zero. At this point, a current control signal (micro-perturbation signal) is generated and superimposed on the current loop input of the servo controller. This micro-perturbation signal can be amplified using a digital adder. Normal servo current control signals sent to the main CPU of flight control Linear superposition is performed to generate the total current control signal. The total current control signal is used to directly drive the power circuit of the subsequent stage, thereby utilizing the inertial compensation force generated by the equivalent mass of the transmission chain. F inertial , offset Figure 3 Preload loss ΔP loss , to achieve F inertial ≥ΔP loss The inertial force compensation effect under certain conditions maintains reliable contact of the force transmission chain at the installation interface.

[0059] Step 4: Implement safety limiting and fault protection measures In embodiments of this application, hard-limiting logic, such as the rated current of the servo motor, can be embedded in the firmware of the servo controller. =5A: Real-time monitoring of the current value corresponding to the total current control signal. If the high-frequency micro-disturbance signal is forcibly set to zero, a normal servo current control signal is output to prevent the servo motor from overheating or saturating due to excessive current. At the same time, a fault flag is set to record abnormal operating conditions and ensure the safe operation of the aircraft's servo control system. If the event of setting the disturbance sinusoidal signal to zero occurs continuously for a preset number of times, such as 3 times, the superposition of the disturbance sinusoidal signal on the servo current control signal is terminated. If the current value corresponding to the total current control signal does not exceed a preset multiple of the rated current, such as 1.1 times, the total current control signal can be directly output to the current loop. After being amplified by the power drive unit, it generates AC power to drive the servo motor, which can maintain reliable contact of the force transmission chain at the installation interface and effectively solve the problem of mechanical clearance and control accuracy deterioration caused by preload drop.

[0060] Figure 4 This is a schematic diagram of the architecture and signal flow of a servo motor flutter stabilization control system according to an embodiment of this application.

[0061] like Figure 4 The diagram illustrates the architecture of a servo motor flutter stabilization control system. The system includes a single-axis accelerometer, a current loop, a power drive unit, a servo motor, a core module area, and a protection and monitoring module 404. The core module area is implemented using an FPGA chip and includes an impact discrimination module 401, a parameter calculation module 402, and a signal synthesis module 403.

[0062] In the embodiments of this application, the output terminal of the single-axis accelerometer is connected to the input terminal of the impact discrimination module 401 to provide the high-frequency acceleration signal of the aircraft in real time.

[0063] The impact discrimination module 401 is connected to the input terminal of the parameter calculation module 402 and outputs a disturbance trigger signal that enters the locked unloading phase.

[0064] The parameter calculation module 402 is used to respond to the received disturbance trigger signal, perform calculations based on the mechanical model to determine the disturbance frequency and disturbance amplitude; and convert the disturbance frequency... The disturbance amplitude A is sent to the signal synthesis module 403. The "flight control command" (i.e., the servo current control signal) can be sent directly to the signal synthesis module 403 from the host computer of the servo controller.

[0065] Signal synthesis module 403 is used for perturbation frequency-based signal synthesis A disturbance signal is generated by the disturbance amplitude A. Upon receiving an enable signal from the protection monitoring module 404, the disturbance signal is superimposed on the servo motor current control signal to obtain a total current control signal. This total current control signal is then sent to the current loop to generate a voltage control signal, which is then converted into a PWM wave by the power drive unit to drive the MOSFET / power transistor in the power topology circuit to output AC power, thereby driving the servo motor. The current loop generates a control signal based on the target current and bus current feedback components corresponding to the total current control signal, so that the power drive unit outputs AC power to drive the servo motor.

[0066] The protection monitoring module 404 is used to collect the bus current feedback signal of the servo motor, generate an enable signal based on the current feedback signal meeting the limiting condition, generate a lockout signal when the current feedback signal does not meet the limiting condition, and send the enable signal or lockout signal to the enable terminal (EN) of the signal synthesis module 403.

[0067] It should be noted that the single-axis accelerometer acquires the aircraft's acceleration signal in real time, providing the basis for unloading phase triggering and perturbation parameter calculation for servo control. The current loop, based on the target current superimposed with the perturbation signal, combined with bus current feedback, achieves precise closed-loop control and overcurrent protection of the servo torque. The power drive unit (such as an H-bridge, MOSFET driver, or intelligent power module IPM) amplifies the signal to AC with sufficient voltage / current to provide the torque and power required for motor rotation.

[0068] In the embodiments of this application, the impact discrimination module 401 can determine whether the impact unloading phase has been entered; firstly, the acceleration signal of the aircraft is collected in real time at a sampling rate of 1 MHz using a single-axis accelerometer, and the acceleration signal is... The sampling time t is fed into the impact discrimination module 401. The impact discrimination module 401 internally executes two criteria simultaneously through a hardware comparator and a differential operation unit. The acceleration value | can be realized through the hardware comparator. | with acceleration threshold The comparison, among which, This can be calculated from the bolt preload at the mounting interface and the equivalent mass of the drive train. If At that time, the change in acceleration can be further determined. If the acceleration value reaches the acceleration threshold and begins to gradually decrease, the system enters the impact unloading phase. When both conditions are met simultaneously, the impact discrimination module 401 outputs a high-level disturbance trigger signal, which is sent to the parameter calculation module 402, indicating that the aircraft's servo control system has entered the impact unloading phase.

[0069] The high-frequency disturbance frequency and amplitude can be calculated using the parameter calculation module 402. After receiving the disturbance trigger signal, the parameter calculation module 402 calculates the disturbance parameters (disturbance frequency) based on a preset mechanical model and a lookup table method. The disturbance amplitude A); the disturbance frequency can be calculated using equation (1) of the mechanical model. :in, For the installation interface contact stiffness, For equivalent mass, k is the safety factor. This is the frequency offset, while also considering automatic avoidance of the structural fundamental frequency and limiting the disturbance frequency within the motor bandwidth. The disturbance amplitude A can be obtained by directly calling the corresponding gear amplitude from the gear lookup table based on the maximum acceleration value corresponding to the current peak impact overload. The parameter calculation module 402 then... A is output in parallel to signal synthesis module 403.

[0070] Signal generation and superposition can be performed through signal synthesis module 403; signal synthesis module 403 internally includes a direct digital synthesis (DDS) submodule and a digital adder: utilizing the perturbation frequency The disturbance amplitude A generates a high-frequency micro-disturbance sinusoidal signal. The disturbance signal is linearly superimposed with the "flight control command" sent by the host computer in an adder to generate a total current control signal; the superimposed total current control signal is output to the current loop as the current loop reference for the servo motor.

[0071] The current loop adjusts the output target current according to the total current control signal, and the AC current, after being amplified by the power drive, drives the servo motor. Since the torque of the servo motor is proportional to the current, the current loop, through bus current feedback, ensures that the motor current equals the target current value, guaranteeing stable servo output torque unaffected by load fluctuations, such as changes in control surface drag or mechanical friction. The torque generated by the motor is transmitted to the control surface through the transmission chain. Simultaneously, the equivalent mass of the transmission chain itself generates inertial force under high-frequency disturbances, offsetting the preload loss at the mounting interface and preventing interface loosening.

[0072] The protection monitoring module 404 can detect the bus current of the servo motor. If the bus current feedback signal exceeds a limit, a lockout process is executed to stop the superposition of high-frequency micro-disturbance signals onto the servo motor current control signal. The protection monitoring module 404 can collect the bus current feedback signal in real time and compare it with the servo motor's rated current threshold. If the current exceeds the limit for three consecutive PWM cycles, the protection monitoring module 404 immediately outputs a lockout signal by pulling down the enable (EN) terminal of the signal synthesis module 403. Upon receiving the lockout signal, the signal synthesis module 403 forcibly resets the high-frequency micro-disturbance signal to zero, retaining only normal flight control commands to prevent motor overheating or damage.

[0073] Figure 5 This is a structural block diagram of a servo motor flutter stabilization control device according to an embodiment of this application.

[0074] like Figure 5 The diagram illustrates a servo motor flutter stabilization control device 500. The device includes, in sequence, an impact discrimination module 501, a parameter calculation module 502, a signal synthesis module 503, and a protection monitoring module 504. Each functional module is integrated based on an FPGA chip, providing millisecond-level response capability. All functional modules are implemented using FPGA IP cores.

[0075] The impact discrimination module 501 is connected to a single-axis accelerometer at its input, with a fixed sampling frequency of 1 MHz. Internally, the impact discrimination module 501 integrates a hardware comparator, a differentiating circuit, and a critical acceleration value calculation unit, enabling real-time determination of whether the impact unloading conditions are met.

[0076] The parameter calculation module 502 receives the disturbance trigger signal from the impact discrimination module and has a pre-set mechanical model and a grading table for quickly calculating the disturbance frequency. The disturbance amplitude A. In this embodiment, the parameter calculation module 502 can be implemented based on the CORDIC (Coordinate Rotation Digital Calculation Method) algorithm hardware acceleration structure, with a calculation latency of less than 10. μs .

[0077] The signal synthesis module 503 integrates a DDS sine wave generator and a digital adder, which can linearly superimpose high-frequency micro-disturbance signals with servo motor current control signals. The superimposed total current control signal is output to the current loop as a current control command. The power drive unit (e.g., H-bridge, MOSFET driver, intelligent power module IPM) can adjust the duty cycle of the PWM wave generated by the internal PWM generation unit to drive the switching MOSFET, generating AC power to drive the servo motor.

[0078] The protection monitoring module 504 collects the bus current feedback signal in real time and compares it with the current threshold. The current threshold is a preset multiple of the rated current of the servo motor, such as 1.2 times or 1.5 times. If the current exceeds the limit for three consecutive PWM cycles, an "over-limit lockout" signal is output to the signal synthesis module 503 to forcibly stop the injection of high-frequency micro-disturbance signals into the normal servo current control signal.

[0079] This application provides a servo motor flutter stabilization control device that can inject micro-disturbances only during the critical window of impact unloading. It can maintain reliable contact of the force transmission chain at the installation interface in milliseconds without additional hardware, effectively suppressing the control dead zone caused by preload drop, improving the control surface following accuracy and the stability of aircraft attitude control, and ensuring the safe operation of the servo motor through a current protection mechanism, making it suitable for the harsh operating conditions of high-maneuverability aircraft.

[0080] This application also provides a servo flutter stabilization control system, the system comprising: a single-axis accelerometer, the servo flutter stabilization control device described in the foregoing embodiments, a current loop, a power drive unit, and a servo motor. The single-axis accelerometer is connected to the input terminal of the servo flutter stabilization control device to provide it with the aircraft's acceleration signal; the output terminal of the servo flutter stabilization control device is connected to the input terminal of the current loop; the output terminal of the current loop is connected to the power drive unit; and the power drive unit drives the servo motor.

[0081] This application proposes a servo motor flutter stabilization control method, device, and system. By superimposing a micro-disturbance sinusoidal signal synchronized with the impact state onto the current loop of the servo motor controller, and when the impact unloading condition is met, the disturbance signal is calculated based on a mechanical model and superimposed onto the servo motor current control signal. The inertial force generated by the equivalent mass of the transmission chain dynamically compensates for the interface pressure loss during the impact unloading stage. This solves the problems of instantaneous drop in preload force at the servo motor mounting interface, loosening of the transmission chain, and deterioration of control accuracy under high overload impact. It achieves millisecond-level reliable contact of the interface transmission chain without additional hardware, effectively suppresses the control dead zone introduced by mechanical backlash, and ensures the servo surface following accuracy and the stability of aircraft attitude control.

[0082] The above are merely preferred embodiments of this application and do not limit the patent scope of this application. Any equivalent structural or procedural transformations made using the content of this application's specification and drawings, or direct or indirect applications in other related technical fields, are similarly included within the patent protection scope of this application.

Claims

1. A servo motor flutter stabilization control method, applied to the servo motor controller of an aircraft, characterized in that, include: When the acceleration signal of the aircraft meets the impact unloading condition, it is determined that the aircraft has entered the impact unloading stage and triggers the output disturbance request signal; the acceleration signal is the acceleration value output by the acceleration sensor acquired at a fixed sampling frequency. In response to the disturbance request signal, the disturbance frequency is calculated based on the mechanical model, and the disturbance amplitude is obtained by retrieving a matching range lookup table based on the maximum acceleration value; the range lookup table is a correspondence between the intervals of acceleration values ​​and the disturbance amplitudes created based on the mechanical equilibrium equations. Based on the disturbance frequency and the disturbance amplitude, a disturbance sinusoidal signal is synthesized; and the disturbance sinusoidal signal is linearly superimposed with the servo motor current control signal to obtain the total current control signal. When the absolute value of the total current control signal is greater than the current threshold, the disturbance sinusoidal signal is set to zero and a fault flag is recorded; the current threshold is a preset multiple of the rated current of the servo motor.

2. The method as described in claim 1, characterized in that, The determination that the aircraft has entered the impact unloading phase when the aircraft's acceleration signal meets the impact unloading conditions includes: If the absolute value of the acceleration signal is greater than the acceleration threshold and the rate of change of the acceleration signal is negative, the impact unloading condition is determined to be met, and the aircraft enters the impact unloading phase. The acceleration threshold is dynamically determined based on the bolt preload and the equivalent mass of the transmission chain at the current installation interface.

3. The method as described in claim 1, characterized in that, The calculation of the disturbance frequency based on the mechanical model includes: The disturbance frequency is calculated based on the following formula; in, For the perturbation frequency, For the contact stiffness of the installation interface, The equivalent mass of the transmission chain. For safety reasons, This is the frequency offset.

4. The method as described in claim 1, characterized in that, The method further includes: The disturbance amplitude A and the maximum acceleration are obtained based on the inverse formula of the mechanical equilibrium equation. The correspondence; in, ξ For safety margin, m load This refers to the total load mass of the control surfaces and transmission mechanism. To achieve the maximum acceleration during the impact unloading phase, The equivalent mass of the transmission chain. For the perturbation frequency, The angular frequency of the disturbance signal; A grading comparison table is established based on the correspondence between the disturbance amplitude and the range of the maximum acceleration.

5. The method as described in claim 1, characterized in that, After setting the disturbance sinusoidal signal to zero when the absolute value of the total current control signal is greater than the current threshold, the method further includes: When the event of setting the disturbance sinusoidal signal to zero occurs continuously for a preset number of times, the superposition of the disturbance sinusoidal signal on the servo current control signal is terminated.

6. A servo motor flutter stabilization control device, integrated within the servo motor controller of an aircraft, characterized in that, The device includes: The impact detection module determines that the aircraft has entered the impact unloading phase and triggers the output disturbance request signal when the acceleration signal of the aircraft meets the impact unloading conditions; the acceleration signal is the acceleration value output by the acceleration sensor acquired at a fixed sampling frequency. The parameter calculation module, in response to the disturbance request signal, calculates the disturbance frequency based on the mechanical model, and obtains the disturbance amplitude based on the maximum acceleration by retrieving a matching range lookup table; the range lookup table is a correspondence between the intervals of acceleration values ​​and the disturbance amplitudes created based on the mechanical equilibrium equations. The signal synthesis module synthesizes a disturbance sinusoidal signal based on the disturbance frequency and the disturbance amplitude; and linearly superimposes the disturbance sinusoidal signal with the servo motor current control signal to obtain the total current control signal. The protection monitoring module sets the disturbance sinusoidal signal to zero and records the fault flag when the absolute value of the total current control signal is greater than the current threshold; the current threshold is a preset multiple of the motor's rated current.

7. The apparatus as claimed in claim 6, characterized in that, The protection and monitoring module is also used to terminate the superposition of the disturbance sinusoidal signal on the servo motor current control signal when the event of the disturbance sinusoidal signal being set to zero occurs continuously for a preset number of times.

8. A servo motor chatter stabilization control system, characterized in that, include: A single-axis accelerometer, a servo motor flutter stabilization control device as described in claim 6 or 7, a current loop, a power drive unit, and a servo motor; The single-axis accelerometer is connected to the input of the servo flutter stabilization control device to provide it with the aircraft's acceleration signal; the output of the servo flutter stabilization control device is connected to the input of the current loop; the output of the current loop is connected to the power drive unit; and the power drive unit is used to drive the servo motor.

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