Aircraft Overload Control Method Based on Angular Velocity Command Generation and Dual Hysteresis Correction

Abstract

This invention relates to the field of aircraft control technology, and particularly to an aircraft overload control method based on angular velocity command generation and dual-hysteresis correction. The method includes: generating an angular velocity command based on an overload command, and feeding back the actual angular velocity to adjust the damping ratio to a preset ideal range; executing a dual-hysteresis series correction strategy to cancel the zeros of the target transfer function, and adjusting the gain to meet the preset high control gain requirements in the low-frequency band, achieving the preset desired open-loop cutoff frequency and phase margin by configuring zeros and poles; and generating a control surface deflection angle command by combining the dual-hysteresis series correction strategy and the damping adjustment loop to control the deflection of the aircraft control surfaces. This solves the problems caused by the fact that related technologies typically use integral control in the overload control loop to reduce steady-state error, which easily introduces phase hysteresis and reduces the system phase margin, leading to slower overload response, increased dynamic overshoot, and even system oscillation.

Description

Technical Field

[0001] This invention relates to the field of aircraft control technology, and in particular to an aircraft overload control method based on angular velocity command generation and dual hysteresis correction. Background Technology

[0002] In related technologies, because the actuators of tail-controlled aircraft are located at the tail, the overload channel often exhibits significant non-minimum phase characteristics, and its system zero point is located in the right half-plane, resulting in an inherent contradiction between rapid response and stability of the controlled object. To reduce steady-state error, related technologies often introduce integral control loops into the overload control loop to improve command tracking accuracy.

[0003] However, in related technologies, integral control loops tend to amplify system phase lag and reduce phase margin in non-minimum phase systems. Especially under conditions of high overload maneuvers or rapid command changes, this can easily lead to problems such as response lag, increased overshoot, or even oscillations, thereby limiting the dynamic performance and robustness of aircraft overload control. It is difficult to balance the requirements of speed and stability, and this issue urgently needs to be addressed. Summary of the Invention

[0004] This invention provides an aircraft overload control method based on angular velocity command generation and dual hysteresis correction, to solve the problem in related technologies where integral control is usually used in the overload control loop to reduce steady-state error, which easily introduces phase hysteresis and reduces the system phase margin, resulting in slower overload response speed, increased dynamic overshoot, or even system oscillation, affecting the stability and dynamic performance of aircraft overload control.

[0005] A first aspect of this invention provides an aircraft overload control method based on angular velocity command generation and dual hysteresis correction, comprising the following steps: acquiring an overload command of the aircraft and generating an angular velocity command based on the overload command; acquiring the actual angular velocity of the aircraft and feeding back the actual angular velocity to a damping adjustment loop to adjust the damping ratio to a preset ideal range; calculating the deviation between the angular velocity command and the actual angular velocity to execute a dual hysteresis series correction strategy to cancel the zeros of the target transfer function, and adjusting the gain to meet the preset high control gain requirements in the low-frequency band, and achieving a preset desired open-loop cutoff frequency and phase margin by configuring zeros and poles; and generating a control surface deflection command of the aircraft by combining the dual hysteresis series correction strategy and the damping adjustment loop to drive the controller to control the deflection of the aircraft control surfaces.

[0006] Through the above technical means, the embodiments of the present invention can adjust the damping ratio to a preset ideal range by feeding back the actual angular velocity, and combine it with a dual-hysteresis series correction strategy to configure zeros and poles to achieve the preset desired open-loop cutoff frequency and phase margin, thereby generating the control surface deflection angle command of the aircraft. This can improve the speed and smoothness of angular velocity and normal overload response while ensuring the stability margin of the system, suppressing dynamic overshoot and oscillation phenomena, and meeting the requirements of high-quality overload flight control of the aircraft.

[0007] Optionally, in one embodiment of the present invention, before canceling the zero point of the target transfer function, the method further includes: constructing a longitudinal short-period dynamic linearization model of the aircraft; and using the longitudinal short-period dynamic linearization model to obtain the transfer function from pitch deflection angle to normal overload and the transfer function from pitch deflection angle to pitch angular velocity.

[0008] Through the above technical means, the embodiments of the present invention can use the transfer function from pitch deflection angle to normal overload and the transfer function from pitch deflection angle to pitch angular velocity to analyze the dynamic characteristics of the overload channel and attitude channel of the aircraft, and then determine the zero point position in the target transfer function. This can be used to compensate for or cancel the inherent zero point of the aircraft, suppress the amplitude-frequency gain spike and phase distortion caused by the zero point, improve the frequency domain characteristics of the system, and provide a basis for subsequent controller zero-pole configuration and stability margin design.

[0009] Optionally, in one embodiment of the present invention, the transfer function from the pitch deflection angle to the normal overload is:

[0010] ,

[0011] in, Indicates the pitch deflection angle. Indicates normal overload, , , , , These represent the static gain of the pitch deflection due to normal overload, the zero-frequency parameter of the transfer function, the natural angular frequency of the longitudinal short-period mode, the zero-damping parameter of the transfer function, and the damping ratio of the longitudinal short-period mode, respectively. This represents the complex frequency domain operator introduced by the Laplace transform;

[0012] The transfer function from the pitch deflection angle to the pitch angular velocity is:

[0013] ,

[0014] in, Indicates pitch angular velocity, , These represent the static gain of pitch angular velocity on rudder deflection and the zero-point time constant of the transfer function, respectively.

[0015] Through the above technical means, the embodiments of the present invention can determine the transfer function from pitch deflection angle to normal overload and the transfer function from pitch deflection angle to pitch angular velocity. These functions can be used to characterize the dynamic relationship between the aerodynamic control surface input and overload response and attitude angular velocity response of the aircraft. Furthermore, they can be used to analyze the frequency domain characteristics and phase characteristics of the aircraft in the overload channel and attitude channel, providing a basis for the construction of the target transfer function, the configuration of controller zeros and poles, and the design of system stability margin.

[0016] Optionally, in one embodiment of the present invention, generating the angular velocity command according to the overload command includes: converting the overload command into the angular velocity command through a pre-built command generator, wherein the command generator is constructed based on the relationship between normal overload and pitch angle motion.

[0017] Through the above technical means, embodiments of the present invention can construct a command generator based on the coupling relationship between normal overload and aircraft pitch angle motion. By processing and mapping the normal overload command, the normal overload requirement is transformed into a corresponding pitch angle velocity command, which enables the aircraft to quickly establish an angle of attack under the action of the overload command, thereby achieving a rapid response to normal overload. At the same time, it avoids the adverse effects of non-minimum phase caused by direct overload closed-loop control, and improves the dynamic performance and stability of the overload control process.

[0018] Optionally, in one embodiment of the present invention, the expression of the instruction generator is:

[0019] ,

[0020] in, Indicates instruction generator, Indicates a step overload instruction. Indicates angular velocity command. , , These represent the static gain of the instruction generator, the zero-point time constant of the instruction generator, and the pole-point time constant of the instruction generator, respectively.

[0021] Through the above technical means, the embodiments of the present invention can determine the structure and parameter configuration of the instruction generator, which can be used to map the normal overload instruction into the angular velocity instruction, thereby achieving a fast and stable response to the overload instruction.

[0022] Optionally, in one embodiment of the present invention, the dual-lag series correction strategy includes a first lag correction element design and a second lag correction element design, wherein the calculation formula for the first lag correction element design is:

[0023] ,

[0024] in, This indicates the first lag correction stage. Represents the time constant at the first zero point. Represents the time constant of the first pole. Indicates the first static gain. This represents the complex frequency domain operator introduced by the Laplace transform;

[0025] The calculation formula for the second lag correction element is as follows:

[0026] ,

[0027] in, This indicates the second lag correction stage. This represents the time constant at the second zero point. This represents the time constant of the second pole. This represents the second static gain.

[0028] Through the above technical means, the embodiments of the present invention can design a first lag correction stage and a second lag correction stage and configure them in series. While ensuring that the system has a high control gain in the low frequency band, it can coordinate and shape the zeros and poles in the target transfer function, gradually compensate for the inherent phase lag of the aircraft overload channel, suppress the gain spike in the amplitude-frequency characteristics, and improve the phase margin and stability margin of the system. Thus, without introducing an integral control stage, it can achieve fast and stable angular velocity and normal overload response.

[0029] Optionally, in one embodiment of the present invention, the control formula of the controller is:

[0030] ,

[0031] in, This indicates the pitch deflection command. Indicates pitch angular velocity, Indicates an overload command. This indicates the angular velocity feedback gain of the damping adjustment loop.

[0032] Through the above technical means, the embodiments of the present invention can determine the structural form and parameter configuration of the controller according to the dynamic characteristics of the aircraft pitch channel, so as to achieve a fast and stable response of pitch angular velocity and normal overload to command input, suppress overshoot and oscillation, and improve the dynamic performance and stability of the overload control system.

[0033] A second aspect of the present invention provides an aircraft overload control device based on angular velocity command generation and dual hysteresis correction, comprising: a generation module for acquiring an overload command of the aircraft and generating an angular velocity command based on the overload command; an inner loop control module for acquiring the actual angular velocity of the aircraft and feeding back the actual angular velocity to a damping adjustment loop to adjust the damping ratio to a preset ideal range; an outer loop control module for calculating the deviation between the angular velocity command and the actual angular velocity to execute a dual hysteresis series correction strategy to cancel the zeros of the target transfer function, and adjusting the gain to meet the preset high control gain requirements in the low-frequency band, and achieving a preset desired open-loop cutoff frequency and phase margin by configuring zeros and poles; and a synthesis module for synthesizing the dual hysteresis series correction strategy and the damping adjustment loop to generate a control surface deflection angle command for the aircraft, thereby driving the controller to control the deflection of the aircraft control surfaces.

[0034] Optionally, in one embodiment of the present invention, the aircraft overload control device based on angular velocity command generation and dual hysteresis correction further includes: a construction module for constructing a longitudinal short-period dynamic linearization model of the aircraft; and a determination module for using the longitudinal short-period dynamic linearization model to obtain the transfer function from pitch deflection angle to normal overload and the transfer function from pitch deflection angle to pitch angular velocity.

[0035] Optionally, in one embodiment of the present invention, the transfer function from the pitch deflection angle to the normal overload is:

[0036] ,

[0037] in, Indicates the pitch deflection angle. Indicates normal overload, , , , , These represent the static gain of the pitch deflection due to normal overload, the zero-frequency parameter of the transfer function, the natural angular frequency of the longitudinal short-period mode, the zero-damping parameter of the transfer function, and the damping ratio of the longitudinal short-period mode, respectively. This represents the complex frequency domain operator introduced by the Laplace transform.

[0038] The transfer function from the pitch deflection angle to the pitch angular velocity is:

[0039] ,

[0040] in, Indicates pitch angular velocity, , These represent the static gain of pitch angular velocity on rudder deflection and the zero-point time constant of the transfer function, respectively.

[0041] Optionally, in one embodiment of the present invention, the generation module includes: a conversion unit, configured to convert the overload command into the angular velocity command through a pre-built command generator, wherein the command generator is constructed based on the relationship between normal overload and pitch angle motion.

[0042] Optionally, in one embodiment of the present invention, the expression of the instruction generator is:

[0043] ,

[0044] in, Indicates instruction generator, Indicates a step overload instruction. Indicates angular velocity command. , , These represent the static gain of the instruction generator, the zero-point time constant of the instruction generator, and the pole-point time constant of the instruction generator, respectively.

[0045] Optionally, in one embodiment of the present invention, the dual-lag series correction strategy includes a first lag correction element design and a second lag correction element design, wherein the calculation formula for the first lag correction element design is:

[0046] ,

[0047] in, This indicates the first lag correction stage. Represents the time constant at the first zero point. Represents the time constant of the first pole. Indicates the first static gain. This represents the complex frequency domain operator introduced by the Laplace transform.

[0048] The calculation formula for the second lag correction element is as follows:

[0049] ,

[0050] in, This indicates the second lag correction stage. This represents the time constant at the second zero point. This represents the time constant of the second pole. This represents the second static gain.

[0051] A third aspect of the present invention provides an electronic device, including: a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the program to implement the aircraft overload control method based on angular velocity command generation and dual hysteresis correction as described in the above embodiments.

[0052] A fourth aspect of the present invention provides a computer-readable storage medium storing a computer program that, when executed by a processor, implements the above-described aircraft overload control method based on angular velocity command generation and dual hysteresis correction.

[0053] A fifth aspect of the present invention provides a computer program product, including a computer program, which, when executed, is used to implement the above-described aircraft overload control method based on angular velocity command generation and dual hysteresis correction.

[0054] Additional aspects and advantages of the invention will be set forth in part in the description which follows, and in part will be obvious from the description, or may be learned by practice of the invention. Attached Figure Description

[0055] The above and / or additional aspects and advantages of the present invention will become apparent and readily understood from the following description of the embodiments taken in conjunction with the accompanying drawings, wherein:

[0056] Figure 1 A flowchart of an aircraft overload control method based on angular velocity command generation and dual hysteresis correction according to an embodiment of the present invention;

[0057] Figure 2 This is a schematic diagram illustrating the principle of an aircraft overload control method based on angular velocity command generation and dual hysteresis correction according to an embodiment of the present invention.

[0058] Figure 3 This is a schematic diagram illustrating the principle of an aircraft overload control method based on angular velocity command generation and dual hysteresis correction, according to another embodiment of the present invention.

[0059] Figure 4 This is a schematic diagram of the simulation results of an embodiment of the aircraft overload control method based on angular velocity command generation and double hysteresis correction according to an embodiment of the present invention.

[0060] Figure 5 A block diagram of an aircraft overload control device based on angular velocity command generation and dual hysteresis correction according to an embodiment of the present invention;

[0061] Figure 6 This is a schematic diagram of the structure of an electronic device according to an embodiment of the present invention.

[0062] Figure label:

[0063] 10-Aircraft overload control device based on angular velocity command generation and dual hysteresis correction; 100-Generation module, 200-Inner loop control module, 300-Outer loop control module, 400-Integrated module; 601-Memory, 602-Processor, 603-Communication interface. Detailed Implementation

[0064] Embodiments of the present invention are described in detail below, examples of which are illustrated in the accompanying drawings, wherein the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The embodiments described below with reference to the accompanying drawings are exemplary and intended to explain the present invention, and should not be construed as limiting the present invention.

[0065] The following describes an embodiment of the aircraft overload control method based on angular velocity command generation and dual hysteresis correction, with reference to the accompanying drawings. Addressing the technical problem mentioned in the background art, where related technologies typically employ integral control in the overload control loop to reduce steady-state error, this method easily introduces phase hysteresis and reduces the system phase margin, leading to slower overload response, increased dynamic overshoot, and even system oscillations, thus affecting the stability and dynamic performance of aircraft overload control, this invention provides an aircraft overload control method based on angular velocity command generation and dual hysteresis correction. In this method, the damping ratio is adjusted to a preset ideal range based on the actual angular velocity, and the deviation between the angular velocity command and the actual angular velocity is calculated to execute a dual hysteresis series correction strategy. Zeros and poles are configured to achieve a preset desired open-loop cutoff frequency and phase margin, generating a control surface deflection angle command. This method employs a two-step closed-loop control structure to adjust the system, offsetting gain spikes caused by inherent zeros in aircraft dynamics, smoothing the system's amplitude-frequency characteristics, and, in conjunction with damping control, significantly improving the system's stability margin. It achieves both rapid and smooth angular velocity and normal overload response while ensuring system stability. Simultaneously, it avoids introducing integral control loops, eliminating the risk of integral saturation and the inherent phase lag of the integrator from the control structure's perspective, thus maintaining good dynamic response characteristics even under conditions of high maneuverability and large command changes. It is particularly suitable for the high-quality overload control requirements of super-maneuverable flight scenarios. This solves the problem that related technologies typically use integral control in overload control loops to reduce steady-state error, which easily introduces phase lag and reduces the system's phase margin, leading to slower overload response speed, increased dynamic overshoot, and even system oscillations, affecting the stability and dynamic performance of aircraft overload control.

[0066] Specifically, Figure 1 This is a flowchart illustrating an aircraft overload control method based on angular velocity command generation and dual hysteresis correction, provided as an embodiment of the present invention.

[0067] like Figure 1 As shown, the aircraft overload control method based on angular velocity command generation and dual hysteresis correction includes the following steps:

[0068] In step S101, the overload command of the aircraft is obtained, and the angular velocity command is generated according to the overload command.

[0069] Among them, the overload command of the aircraft can describe the magnitude of the overload that the aircraft is expected to achieve during the maneuver; the angular velocity command can indicate the target angular motion state that the aircraft should achieve, and can include a steady-state component for providing the desired line-of-sight angular velocity and a dynamic component for quickly establishing the airflow angle.

[0070] It can be explained that the overload command for the aircraft can be obtained by the aircraft guidance system calculating and generating it based on flight mission requirements and the current flight status; or it can be determined by the maneuver planning module and flight control law based on the target trajectory, attitude change requirements, and safety constraints; or it can be obtained by mapping or transforming the overload command based on the dynamic relationship between the aircraft's overload and attitude motion to obtain the corresponding angular velocity command. The acquisition and generation methods can be set by those skilled in the art according to the actual situation, and no specific restrictions are imposed here.

[0071] Optionally, in one embodiment of the present invention, generating an angular velocity command based on an overload command includes: converting the overload command into an angular velocity command using a pre-built command generator, wherein the command generator is constructed based on the relationship between normal overload and pitch angle motion.

[0072] It is understandable that, in order to quickly generate normal overload, the embodiments of the present invention can rapidly establish the angle of attack, and the establishment of the angle of attack depends on the rapid response of the aircraft's pitch angle motion. Therefore, the embodiments of the present invention can use an angular velocity command generator to generate a corresponding angular velocity command based on the normal overload command, and use it as a reference input for the angular velocity control loop, so that the aircraft can quickly generate the required pitch angular velocity, thereby accelerating the angle of attack establishment process and achieving a rapid response to normal overload.

[0073] As one possible implementation method, embodiments of the present invention can establish a command generator based on the principle of advance compensation, which can compensate for phase lag in the overload channel and improve the dynamic response performance of the aircraft overload control.

[0074] Optionally, in one embodiment of the present invention, the expression of the instruction generator can be represented as:

[0075] ,

[0076] in, Indicates instruction generator, Indicates a step overload instruction. Indicates angular velocity command. , , These represent the static gain of the instruction generator, the zero-point time constant of the instruction generator, and the pole-point time constant of the instruction generator, respectively. This represents the complex frequency domain operator introduced by the Laplace transform.

[0077] It can be explained that, ;because Zero point is generally located in the low to mid-frequency range, causing the dynamic response from angular velocity to overload to be significantly affected. The influence of zero point, therefore, embodiments of the present invention can select equal The time constant at zero point , It can be designed according to the desired overload dynamic characteristics. When a step overload command is input... At that time, the output Initially, it manifests as a large dynamic pulse, which can be used to drive the control surfaces to deflect rapidly in order to accumulate the angle of attack; as time progresses, the command decays to a steady-state value. This value corresponds to the steady-state line-of-sight angular velocity required to maintain this overload, therefore it can be... Designed as follows: ,in The input compensation gain is added to ensure that the DC gain for angular velocity control is 1.

[0078] In step S102, the actual angular velocity of the aircraft is collected and fed back to the damping adjustment circuit to adjust the damping ratio to a preset ideal range.

[0079] It can be explained that the damping adjustment loop can feed the angular velocity signal back to the forward channel through the inner loop feedback gain, forming a generalized controlled object with the short-cycle dynamics of the aircraft, and then adjust the inner loop feedback gain so that the damping ratio of the generalized controlled object is within the preset ideal damping range.

[0080] In some cases, embodiments of the present invention can acquire the actual angular velocity of the aircraft using angular velocity sensors or inertial measurement units (such as gyroscopes). The damping ratio can describe the oscillation characteristics of the angular velocity or overload response, reflecting how quickly the oscillation amplitude decays over time during the system's response after being disturbed or given a command input.

[0081] In the embodiments of the present invention, the damping ratio within the preset ideal range can be 0.6 to 0.8. Generally, the larger the damping ratio, the faster the system oscillation decays and the smaller the overshoot; if the damping ratio is too small, obvious oscillations are likely to occur. It can be set by those skilled in the art according to the actual situation, and no specific limitation is made here.

[0082] Specifically, embodiments of the present invention can measure the angular velocity of an aircraft in real time and design an internal closed loop for angular velocity control to control damping characteristics; embodiments of the present invention can introduce an inner loop angular velocity feedback gain. At this time, the servo command It can be represented as:

[0083] ,

[0084] in, The equivalent pitch angular velocity command generated by the external closed loop of angular velocity control.

[0085] As can be seen, the embodiments of the present invention can be adjusted. By analyzing the values ​​and closed-loop pole distribution, the damping ratio of the generalized controlled object (aircraft dynamics + inner loop feedback) is made to reach the ideal range of 0.6 to 0.8, which can eliminate the violent oscillations of short-period motion.

[0086] As one possible approach, if the dynamics of the servo motor are ignored, This can be obtained by solving the following quadratic equation (taking the negative root):

[0087] ,

[0088] in, The desired natural angular frequency.

[0089] In step S103, the deviation between the angular velocity command and the actual angular velocity is calculated to execute a double-hysteresis series correction strategy to cancel the zeros of the target transfer function, and the gain is adjusted to meet the preset high control gain requirements of the low-frequency band. The preset desired open-loop cutoff frequency and phase margin are achieved by configuring the zeros and poles.

[0090] It can be explained that the dual-lag series compensation strategy refers to setting two lag compensation elements in series in the control loop. By jointly adjusting the amplitude-frequency and phase-frequency characteristics of the system, the system maintains sufficient steady-state accuracy in the low-frequency range and suppresses noise and adverse dynamic effects in the mid-to-high frequency range, thereby improving the system's stability and dynamic performance. In this embodiment of the invention, the dual-lag series compensation strategy can be used to design an external closed loop for the angular velocity control loop to compensate for or eliminate zeros in the target transfer function, thus reducing the adverse effects of non-minimum phase zeros on the system's dynamic performance.

[0091] In embodiments of the present invention, the preset low-frequency band can refer to the frequency range close to zero in the frequency domain characteristics of the control system, such as 0 to 1 rad / s, used to characterize the steady-state and low-frequency tracking performance of the system. Zeros primarily affect the phase characteristics and transient response of the system; proper zero configuration can provide phase lead, accelerate the response, and improve the phase margin. Poles primarily affect the stability and dynamic speed of the system; the location of poles can determine whether the system response is fast or slow and whether it is prone to oscillation. The open-loop cutoff frequency characterizes the system's response speed and bandwidth; the higher the cutoff frequency, the faster the system response. Phase margin characterizes the "safety margin" of the system from instability; the larger the phase margin, the less prone the system is to oscillation.

[0092] The embodiments of the present invention can make the system respond quickly enough and remain stable by reasonably configuring the zero point and pole, and on this basis, calculate the control deflection angle command to drive the control surface of the aircraft.

[0093] Optionally, in one embodiment of the present invention, before canceling the zero point of the target transfer function, the method further includes: constructing a longitudinal short-period dynamic linearization model of the aircraft; and using the longitudinal short-period dynamic linearization model to obtain the transfer function from pitch deflection angle to normal overload and the transfer function from pitch deflection angle to pitch angular velocity.

[0094] As a specific example, in an embodiment of the present invention, the state variable can be set as the angle of attack. and pitch angular velocity The control variable is the pitch deflection angle. The output is a normal overload. A linearized model of the longitudinal short-period dynamics of the aircraft was constructed.

[0095] Next, from the longitudinal short-period dynamic linearization model, the transfer function from pitch deflection angle to normal overload and the transfer function from pitch deflection angle to pitch angular velocity can be obtained in the embodiments of the present invention.

[0096] Optionally, in one embodiment of the present invention, the transfer function from pitch deflection to normal overload can be expressed as:

[0097] ,

[0098] in, Indicates the pitch deflection angle. Indicates normal overload, , , , , The values ​​represent the static gain of the pitch deflection angle due to normal overload, the zero-point frequency parameter of the transfer function, the natural angular frequency of the longitudinal short-period mode, the zero-point damping parameter of the transfer function, and the damping ratio of the longitudinal short-period mode.

[0099] The transfer function from pitch deflection angle to pitch angular velocity can be expressed as:

[0100] ,

[0101] in, Indicates pitch angular velocity, , This represents the static gain of pitch rate on rudder deflection and the zero-point time constant of the transfer function.

[0102] By analyzing the above formula, we can find that: The presence of zeros in the right half of the complex plane can demonstrate the non-minimum phase characteristic. There exists a true zero point This may cause the open-loop logarithmic amplitude-frequency characteristic curve of the angular velocity control loop to exhibit an upward tilt in amplitude, which is not conducive to system correction.

[0103] Furthermore, based on the above analysis, embodiments of the present invention can design a dual-hysteresis series correction strategy to achieve both speed and stability in the aircraft's normal overload response.

[0104] Optionally, in one embodiment of the present invention, the dual-lag series compensation strategy includes a first lag compensation element design and a second lag compensation element design, wherein the calculation formula for the first lag compensation element design can be expressed as:

[0105] ,

[0106] in, This indicates the first lag correction stage. Represents the time constant at the first zero point. Represents the time constant of the first pole. This represents the first static gain.

[0107] The embodiments of the present invention can make , By canceling out the zeros of the controlled object, the amplitude-frequency curve that was originally raised due to the zeros can be lowered, making the design at the frequency crossing point easier. It can be designed at 5 to 10 times the intended crossing frequency (or it can be designed to be 0, in which case it is a first-order inertial element).

[0108] The calculation formula for the second lag compensation stage can be expressed as:

[0109] ,

[0110] in, This indicates the second lag correction stage. This represents the time constant at the second zero point. This represents the time constant of the second pole. This represents the second static gain.

[0111] It can be explained that, The zero-pole ratio can be determined based on the steady-state characteristics of angular velocity tracking. The time constant can be determined based on the amplitude characteristics at the designed crossover frequency; by designing the zero point to be 1 / 10 to 1 / 2 of the crossover frequency, the time constant can be determined. Pole time constant According to Confirmed; the benefits of this design include, but are not limited to:

[0112] (a) In the low-frequency band, high gain can be provided to ensure steady-state accuracy of angular velocity tracking. Although there is no integrator (the theoretical steady-state error is not zero), the high gain can control the error within the allowable range of engineering.

[0113] (b) In the mid-to-high frequency band, the hysteresis correction element can be used to reduce the crossover frequency, ensure the phase margin, and filter out high-frequency noise and aircraft elastic vibration components in the angular velocity gyroscope measurement.

[0114] In an embodiment of the present invention, the first hysteresis compensation stage can be used to specifically offset the inherent zeros in the angular velocity transfer function, eliminating the upward tilting of the amplitude-frequency characteristic curve in the mid-frequency band caused by them; the second hysteresis compensation stage meets the high control gain requirements in the low-frequency band by adjusting the gain, and achieves the desired open-loop cutoff frequency and phase margin by configuring zeros and poles, thereby ensuring control accuracy without an integral stage.

[0115] In step S104, the dual-hysteresis series correction strategy and the damping adjustment loop are combined to generate the control surface deflection angle command of the aircraft control surface, so as to drive the controller to control the deflection of the aircraft control surface.

[0116] In an embodiment of the present invention, after the above-described internal and external closed loops, the DC gain of the closed-loop transfer function of the angular velocity control loop can be expressed as:

[0117] ,

[0118] General order Therefore, the input compensation gain of the angular velocity control loop can be Thus, the static gain of the instruction generator is obtained. .

[0119] Optionally, in one embodiment of the present invention, the control formula of the controller can be expressed as:

[0120] ,

[0121] in, Indicates the pitch deflection angle. Indicates pitch angular velocity, Indicates a step overload instruction. This indicates the angular velocity feedback gain of the damping adjustment loop.

[0122] like Figure 2 As shown, in an embodiment of the present invention, the controller receives an overload command, which is then processed by the aforementioned command generator. It can calculate the outer loop control command. Subtract the measurement feedback Then, they passed through in sequence. , The inner loop control command is calculated. Combined with the inner ring angular velocity feedback gain The final output is the rudder deflection command. The servo controller drives the control surface to deflect, in order to adjust... , These parameters, along with flight status parameters, enable the aircraft to achieve high-quality overload flight.

[0123] Combination Figure 3 , Figure 4 The following specific example will further illustrate the aircraft overload control method based on angular velocity command generation and dual hysteresis correction according to an embodiment of the present invention.

[0124] like Figure 3 As shown, this embodiment of the invention can convert the overload command of the guidance loop into an angular velocity command containing dynamic components and steady-state line-of-sight angular velocity components through a command generator based on the principle of lead compensation. This command is then input into a dual-closed-loop control loop without an integral component. The inner loop of the control loop establishes ideal damping through angular velocity feedback; the outer loop employs a dual-hysteresis series compensation strategy. The first hysteresis compensation stage cancels the inherent zero of the angular velocity transfer function to suppress amplitude-frequency rise, while the second hysteresis compensation stage provides high gain while ensuring sufficient phase margin, thus obtaining the final control deflection angle command. This command acts on the aircraft's aerodynamic control surfaces to generate corresponding aerodynamic forces and moments, driving changes in the aircraft's attitude and normal overload through the aircraft's dynamic processes.

[0125] like Figure 4As shown, the simulation results of the aircraft overload control method based on angular velocity command generation and double hysteresis correction in this embodiment of the invention can be obtained by taking normal overload control as an example. It can be seen that after the overload command undergoes a step change, the normal overload response can quickly follow the command change. Although there is a slight transient reverse change in the initial stage, the response process is stable, the overshoot is small, and it can converge to the command value in a short time. This shows that the aircraft overload control based on angular velocity command generation and double hysteresis correction in this embodiment of the invention can achieve rapid establishment of normal overload while ensuring system stability, thus verifying the effectiveness of the method in improving dynamic response performance.

[0126] The aircraft overload control method based on angular velocity command generation and dual-hysteresis correction proposed in this invention adjusts the damping ratio to a preset ideal range by feeding back the actual angular velocity, and combines it with a dual-hysteresis series correction strategy to configure zeros and poles to achieve the preset desired open-loop cutoff frequency and phase margin, thereby generating the control surface deflection angle command. This method can employ a two-step closed-loop control structure to adjust the system, offsetting the gain spikes caused by the inherent zeros in aircraft dynamics, smoothing the system's amplitude-frequency characteristics, and, in conjunction with damping control, significantly improving the system's stability margin. It can achieve both rapid and stable angular velocity and normal overload responses while ensuring system stability. Simultaneously, it avoids introducing integral control loops, eliminating the risk of integral saturation and the adverse effects of the integrator's inherent phase hysteresis on system stability from the control structure perspective. Therefore, it can maintain good dynamic response characteristics even under conditions of high maneuverability and large command changes, and is particularly suitable for high-quality overload control requirements in hyper-maneuverable flight scenarios.

[0127] Next, referring to the accompanying drawings, we describe the aircraft overload control device based on angular velocity command generation and dual hysteresis correction proposed according to an embodiment of the present invention.

[0128] Figure 5 This is a block diagram of an aircraft overload control device based on angular velocity command generation and dual hysteresis correction according to an embodiment of the present invention.

[0129] like Figure 5 As shown, the aircraft overload control device 10 based on angular velocity command generation and dual hysteresis correction includes: a generation module 100, an inner loop control module 200, an outer loop control module 300, and an integrated module 400.

[0130] The generation module 100 is used to obtain the overload command of the aircraft and generate the angular velocity command based on the overload command.

[0131] The inner loop control module 200 is used to collect the actual angular velocity of the aircraft and provide feedback, and then feed the actual angular velocity back to the damping adjustment loop to adjust the damping ratio to a preset ideal range.

[0132] The external loop control module 300 is used to calculate the deviation between the angular velocity command and the actual angular velocity in order to execute a double-hysteresis series correction strategy to cancel the zeros of the target transfer function and adjust the gain to meet the high control gain requirements of the preset low frequency band. By configuring the zeros and poles, the preset desired open-loop cutoff frequency and phase margin can be achieved.

[0133] The integrated module 400 integrates a dual-hysteresis series correction strategy and a damping adjustment loop to generate control deflection angle commands for the aircraft's control surfaces, thereby driving the controller to control the deflection of the aircraft's control surfaces.

[0134] Optionally, in one embodiment of the present invention, the aircraft overload control device 10 based on angular velocity command generation and dual hysteresis correction further includes: a construction module and a determination module.

[0135] Among them, the construction module is used to construct a linearized longitudinal short-period dynamics model of the aircraft;

[0136] The determination module is used to obtain the transfer function from pitch deflection to normal overload and the transfer function from pitch deflection to pitch angular velocity using the longitudinal short-period dynamic linearization model.

[0137] Optionally, in one embodiment of the present invention, the transfer function from pitch deflection to normal overload is:

[0138] ,

[0139] in, Indicates the pitch deflection angle. Indicates normal overload, , , , , The values ​​represent the static gain of the pitch deflection angle due to normal overload, the zero-point frequency parameter of the transfer function, the natural angular frequency of the longitudinal short-period mode, the zero-point damping parameter of the transfer function, and the damping ratio of the longitudinal short-period mode.

[0140] The transfer function from pitch deflection angle to pitch angular velocity is:

[0141] ,

[0142] in, Indicates pitch angular velocity, , This represents the static gain of pitch rate on rudder deflection and the zero-point time constant of the transfer function.

[0143] Optionally, in one embodiment of the present invention, the generation module 100 includes a conversion unit.

[0144] The conversion unit is used to convert overload commands into angular velocity commands through a pre-built command generator, wherein the command generator is constructed based on the relationship between normal overload and pitch angle motion.

[0145] Optionally, in one embodiment of the present invention, the expression of the instruction generator is:

[0146] ,

[0147] in, Indicates instruction generator, Indicates a step overload instruction. Indicates angular velocity command. , , This represents the static gain of the instruction generator, the zero-point time constant of the instruction generator, and the pole-point time constant of the instruction generator.

[0148] Optionally, in one embodiment of the present invention, the dual-hysteresis series correction strategy includes a first hysteresis correction element design and a second hysteresis correction element design, wherein,

[0149] The calculation formula for the first lag correction stage is as follows:

[0150] ,

[0151] in, This indicates the first lag correction stage. Represents the time constant at the first zero point. Represents the time constant of the first pole. Indicates the first static gain. This represents the complex frequency domain operator introduced by the Laplace transform.

[0152] The calculation formula for the second lag correction element is as follows:

[0153] ,

[0154] in, This indicates the second lag correction stage. This represents the time constant at the second zero point. This represents the time constant of the second pole. This represents the second static gain.

[0155] Optionally, in one embodiment of the present invention, the control formula of the controller is:

[0156] ,

[0157] in, This indicates the pitch deflection command. Indicates pitch angular velocity, Indicates an overload command. This indicates the angular velocity feedback gain of the damping adjustment loop.

[0158] It should be noted that the foregoing explanation of the aircraft overload control method embodiment based on angular velocity command generation and dual hysteresis correction also applies to the aircraft overload control device based on angular velocity command generation and dual hysteresis correction in this embodiment, and will not be repeated here.

[0159] The aircraft overload control device based on angular velocity command generation and dual hysteresis correction proposed in this invention adjusts the damping ratio to a preset ideal range by feeding back the actual angular velocity, and combines a dual hysteresis series correction strategy to configure zeros and poles to achieve the preset desired open-loop cutoff frequency and phase margin, thereby generating the control surface deflection angle command. A two-step closed-loop control structure can be used to adjust the system to offset the gain spike caused by the inherent zeros of the aircraft dynamics, smooth the system's amplitude-frequency characteristics, and, in conjunction with damping control, significantly improve the system's stability margin. It can achieve rapid and smooth angular velocity and normal overload response while ensuring system stability. At the same time, it can avoid introducing integral control loops, eliminating the risk of integral saturation and the adverse effects of the inherent phase hysteresis of the integrator on system stability from the control structure. Thus, it can maintain good dynamic response characteristics even under conditions of large maneuvers and large command changes, and is particularly suitable for the high-quality overload control requirements in super-maneuverable flight scenarios.

[0160] Figure 6 This is a schematic diagram of an electronic device provided in an embodiment of the present invention. The electronic device may include:

[0161] The memory 601, the processor 602, and the computer program stored on the memory 601 and capable of running on the processor 602.

[0162] When the processor 602 executes the program, it implements the aircraft overload control method based on angular velocity command generation and dual hysteresis correction provided in the above embodiments.

[0163] Furthermore, electronic devices also include:

[0164] Communication interface 603 is used for communication between memory 601 and processor 602.

[0165] The memory 601 is used to store computer programs that can run on the processor 602.

[0166] The memory 601 may include high-speed RAM memory, and may also include non-volatile memory, such as at least one disk storage device.

[0167] If the memory 601, processor 602, and communication interface 603 are implemented independently, then the communication interface 603, memory 601, and processor 602 can be interconnected via a bus to complete communication between them. The bus can be an Industry Standard Architecture (ISA) bus, a Peripheral Component Interconnect (PCI) bus, or an Extended Industry Standard Architecture (EISA) bus, etc. Buses can be categorized as address buses, data buses, control buses, etc. For ease of representation, Figure 6 The bus is represented by a single thick line, but this does not mean that there is only one bus or one type of bus.

[0168] Optionally, in a specific implementation, if the memory 601, processor 602, and communication interface 603 are integrated on a single chip, then the memory 601, processor 602, and communication interface 603 can communicate with each other through an internal interface.

[0169] Processor 602 may be a central processing unit (CPU), an application specific integrated circuit (ASIC), or one or more integrated circuits configured to implement embodiments of the present invention.

[0170] This embodiment also provides a computer-readable storage medium storing a computer program that, when executed by a processor, implements the above-described aircraft overload control method based on angular velocity command generation and dual hysteresis correction.

[0171] This invention also provides a computer program product, including a computer program that can run computer instructions. When these computer instructions are executed by a processor, they implement the aircraft overload control method based on angular velocity instruction generation and dual hysteresis correction provided in this invention.

[0172] In the description of this specification, the references to terms such as "one embodiment," "some embodiments," "example," "specific example," or "some examples," etc., indicate that a specific feature, structure, material, or characteristic described in connection with that embodiment or example is included in at least one embodiment or example of the present invention. In this specification, the illustrative expressions of the above terms do not necessarily refer to the same embodiment or example. Furthermore, the specific features, structures, materials, or characteristics described may be combined in any suitable manner in one or more embodiments or examples. Moreover, without contradiction, those skilled in the art can combine and integrate the different embodiments or examples described in this specification, as well as the features of different embodiments or examples.

[0173] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of indicated technical features. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one of that feature. In the description of this invention, "N" means at least two, such as two, three, etc., unless otherwise explicitly specified.

[0174] Any process or method description in the flowchart or otherwise herein can be understood as representing a module, segment, or portion of code comprising one or N executable instructions for implementing custom logic functions or processes, and the scope of preferred embodiments of the invention includes additional implementations in which functions may be performed not in the order shown or discussed, including substantially simultaneously or in reverse order depending on the functions involved, as will be understood by those skilled in the art to which embodiments of the invention pertain.

[0175] The logic and / or steps represented in the flowchart or otherwise described herein, for example, can be considered as a sequenced list of executable instructions for implementing logical functions, and can be embodied in any computer-readable medium for use by, or in conjunction with, an instruction execution system, apparatus, or device (such as a computer-based system, a processor-included system, or other system that can fetch and execute instructions from, an instruction execution system, apparatus, or device). For the purposes of this specification, "computer-readable medium" can be any means that can contain, store, communicate, propagate, or transmit programs for use by, or in conjunction with, an instruction execution system, apparatus, or device. More specific examples (a non-exhaustive list) of computer-readable media include: an electrical connection having one or more wires (electronic device), a portable computer disk drive (magnetic device), random access memory (RAM), read-only memory (ROM), erasable and editable read-only memory (EPROM or flash memory), fiber optic devices, and portable optical disc read-only memory (CDROM). Alternatively, the computer-readable medium may be paper or other suitable media on which the program can be printed, since the program can be obtained electronically by optically scanning the paper or other medium, followed by editing, interpreting, or otherwise processing as necessary, and then stored in a computer memory.

[0176] It should be understood that various parts of the present invention can be implemented in hardware, software, firmware, or a combination thereof. In the above embodiments, the N steps or methods can be implemented in software or firmware stored in memory and executed by a suitable instruction execution system. If implemented in hardware, as in another embodiment, it can be implemented using any one or more of the following techniques known in the art: discrete logic circuits having logic gates for implementing logical functions on data signals, application-specific integrated circuits (ASICs) having suitable combinational logic gates, programmable gate arrays (PGAs), field-programmable gate arrays (FPGAs), etc.

[0177] Those skilled in the art will understand that all or part of the steps of the methods in the above embodiments can be implemented by a program instructing related hardware. The program can be stored in a computer-readable storage medium, and when executed, the program includes one or a combination of the steps of the method embodiments.

[0178] Furthermore, the functional units in the various embodiments of the present invention can be integrated into a processing module, or each unit can exist physically separately, or two or more units can be integrated into a module. The integrated module can be implemented in hardware or as a software functional module. If the integrated module is implemented as a software functional module and sold or used as an independent product, it can also be stored in a computer-readable storage medium.

[0179] The storage medium mentioned above can be a read-only memory, a disk, or an optical disk, etc. Although embodiments of the present invention have been shown and described above, it is to be understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those skilled in the art can make changes, modifications, substitutions, and variations to the above embodiments within the scope of the present invention.

Claims

1. An aircraft overload control method based on angular velocity command generation and dual hysteresis correction, characterized in that, Includes the following steps: Obtain the overload command of the aircraft and generate an angular velocity command based on the overload command; The actual angular velocity of the aircraft is collected and fed back to the damping adjustment circuit to adjust the damping ratio to a preset ideal range; The deviation between the angular velocity command and the actual angular velocity is calculated to execute a double-hysteresis series correction strategy to cancel the zeros of the target transfer function, and the gain is adjusted to meet the high control gain requirements of the preset low-frequency band. The preset desired open-loop cutoff frequency and phase margin are achieved by configuring the zeros and poles. The combined dual-hysteresis series correction strategy and the damping adjustment loop generate the deflection angle command of the aircraft control surface, thereby driving the controller to control the deflection of the aircraft control surface; The dual-hysteresis series compensation strategy includes the design of a first hysteresis compensation element and a second hysteresis compensation element, wherein... The calculation formula for the first hysteresis compensation element is as follows: ， in, This indicates the first lag correction stage. Represents the time constant at the first zero point. Represents the time constant of the first pole. Indicates the first static gain. This represents the complex frequency domain operator introduced by the Laplace transform; The calculation formula for the second lag correction element is as follows: ， in, This indicates the second lag correction stage. This represents the time constant at the second zero point. This represents the time constant of the second pole. This represents the second static gain.

2. The aircraft overload control method based on angular velocity command generation and dual hysteresis correction according to claim 1, characterized in that, Before canceling the zeros of the target transfer function, the method further includes: Construct a longitudinal short-period dynamic linearization model of the aircraft; The transfer functions from pitch deflection angle to normal overload and from pitch deflection angle to pitch angular velocity are obtained using the aforementioned longitudinal short-period dynamic linearization model.

3. The aircraft overload control method based on angular velocity command generation and dual hysteresis correction according to claim 2, characterized in that, in, The transfer function from the pitch deflection angle to the normal overload is: ， in, Indicates the pitch deflection angle. Indicates normal overload, , , , , These represent the static gain of the pitch deflection due to normal overload, the zero-frequency parameter of the transfer function, the natural angular frequency of the longitudinal short-period mode, the zero-damping parameter of the transfer function, and the damping ratio of the longitudinal short-period mode, respectively. This represents the complex frequency domain operator introduced by the Laplace transform; The transfer function from the pitch deflection angle to the pitch angular velocity is: ， in, Indicates pitch angular velocity, , These represent the static gain of pitch angular velocity on rudder deflection and the zero-point time constant of the transfer function, respectively.

4. The aircraft overload control method based on angular velocity command generation and dual hysteresis correction according to claim 1, characterized in that, The step of generating angular velocity commands based on the overload commands includes: The overload command is converted into the angular velocity command by a pre-built command generator, wherein the command generator is constructed based on the relationship between normal overload and pitch angle motion.

5. The aircraft overload control method based on angular velocity command generation and dual hysteresis correction according to claim 4, characterized in that, The expression for the instruction generator is: ， in, This refers to the instruction generator. Indicates a step overload instruction. Indicates angular velocity command, , , These represent the static gain of the instruction generator, the zero-point time constant of the instruction generator, and the pole-point time constant of the instruction generator, respectively.

6. The method according to claim 1, characterized in that, The control formula for the controller is: ， in, This indicates the pitch deflection command. Indicates pitch angular velocity, Indicates an overload command. This indicates the angular velocity feedback gain of the damping adjustment loop. , , These represent the static gain of the instruction generator, the zero-point time constant of the instruction generator, and the pole-point time constant of the instruction generator, respectively.

7. An aircraft overload control device based on angular velocity command generation and dual hysteresis correction, characterized in that, include: The generation module is used to obtain the overload command of the aircraft and generate an angular velocity command based on the overload command; The inner loop control module is used to collect the actual angular velocity of the aircraft and feed the actual angular velocity back to the damping adjustment loop to adjust the damping ratio to a preset ideal range; The external loop control module is used to calculate the deviation between the angular velocity command and the actual angular velocity in order to execute a double-hysteresis series correction strategy to cancel the zeros of the target transfer function and adjust the gain to meet the high control gain requirements of the preset low frequency band. By configuring the zeros and poles, the preset desired open-loop cutoff frequency and phase margin can be achieved. The integrated module combines the dual-hysteresis series correction strategy and the damping adjustment loop to generate the control surface deflection command of the aircraft, so as to drive the controller to control the deflection of the aircraft control surface; The dual-hysteresis series compensation strategy includes the design of a first hysteresis compensation element and a second hysteresis compensation element, wherein... The calculation formula for the first hysteresis compensation element is as follows: ， in, This indicates the first lag correction stage. Represents the time constant at the first zero point. Represents the time constant of the first pole. Indicates the first static gain. This represents the complex frequency domain operator introduced by the Laplace transform; The calculation formula for the second lag correction element is as follows: ， in, This indicates the second lag correction stage. This represents the time constant at the second zero point. This represents the time constant of the second pole. This represents the second static gain.

8. An electronic device, characterized in that, include: The system includes a memory, a processor, and a computer program stored in the memory and capable of running on the processor, wherein the processor executes the program to implement the aircraft overload control method based on angular velocity command generation and dual hysteresis correction as described in any one of claims 1-6.

9. A computer-readable storage medium having a computer program stored thereon, characterized in that, The program is executed by the processor to implement the aircraft overload control method based on angular velocity command generation and dual hysteresis correction as described in any one of claims 1-6.

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