Simulation method and device for aileron fault of medium-sized propeller-driven aircraft, computer equipment and medium
By setting aileron failure modes, establishing control surface actuator models, and correcting aerodynamic parameters, the problem of coarse aileron simulation models in existing technologies has been solved, achieving more accurate aircraft aileron failure simulation and improving the predictive capability for flight safety.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-11-26
- Publication Date
- 2026-04-14
AI Technical Summary
Existing simulation models for aircraft aileron failures are crude, with poor consistency between theoretical simulations and real aircraft models, and fail to effectively simulate the impact of aileron failures on flight safety.
The aileron failure mode is set, the control surface actuator model is established, the six-component aerodynamic parameters are obtained, the propeller slipstream asymmetric correction is performed, the aircraft state parameters are simulated using the six-degree-of-freedom simulation model, the dynamic mapping table is introduced to adjust the oscillation parameters, and the slipstream effect is considered.
It improves the accuracy and consistency of the aileron failure simulation model, enhances the predictive ability for flight safety, provides a more comprehensive basis for emergency response, and is suitable for engineering real-time simulation systems.
Smart Images
Figure CN121859522A_ABST
Abstract
Description
Technical Field
[0001] This invention relates to the field of simulation technology, and in particular to a simulation method, apparatus, computer equipment, and medium for aileron failure of a medium-sized propeller aircraft. Background Technology
[0002] With the booming development of the aviation industry, flight safety has always been a focus of attention. As a key component for controlling the attitude and stability of an aircraft, the reliability of its performance is directly related to flight safety.
[0003] As one of the primary control surfaces of an aircraft, the aileron is responsible for roll attitude control, directly affecting flight safety. However, in actual use, the aileron can malfunction for various reasons, such as jamming, damage, and control failure, which can lead to serious flight accidents. Therefore, effectively predicting the flight response after aileron malfunction is of great significance for ensuring flight safety.
[0004] In recent years, to predict the impact of various potential control surface failures on flight safety, flight simulation technology has been widely used to simulate aircraft flight responses under various operating conditions, providing a theoretical basis for the compilation of flight manuals. By establishing a six-degree-of-freedom motion simulation model of the aircraft and adding various aileron failures, it is possible to predict the flight response after an aileron failure and formulate emergency response procedures. However, existing simulation methods still have certain limitations when dealing with aileron failures, such as ignoring the response delay of the flight control system, not using actuator models or overly simplifying actuator models, and having a single control surface failure mode. Summary of the Invention
[0005] In view of this, embodiments of the present invention provide a simulation method for aileron failure in a medium-sized propeller aircraft, to solve the technical problems of coarse simulation models and poor consistency between theoretical simulations and real aircraft models in existing technologies for aileron failures. The method includes: Set the failure modes of the aileron, and set the deflection rate ratio and control surface target deflection for each failure mode of the aileron. The failure modes of the aileron include jamming, half-speed deflection, return to center, sharp deflection, damped and non-command oscillation. Establish a control surface actuator model, input the control surface hinge torque, the current control surface deflection, the target control surface deflection, and the deflection rate ratio into the control surface actuator model, and output the control surface deflection of each aileron through the control surface actuator model; Obtain the six-component aerodynamic parameters of a single aileron. If these parameters cannot be obtained, estimate the six-component aerodynamic parameters of a single aileron using the three-component aerodynamic parameters of the aileron as a whole. The six-component aerodynamic parameters are modified by propeller slipflow asymmetry to generate asymmetric six-component aerodynamic parameters; Based on the control surface deflection of each aileron and the asymmetric six-component aerodynamic parameters, the state parameters of the aircraft at the next moment are obtained by simulating the state parameters of the aircraft at the previous moment using the six-degree-of-freedom simulation model.
[0006] This invention also provides a simulation device for aileron failure in a medium-sized propeller aircraft, addressing the technical problems of coarse simulation models and poor consistency between theoretical simulations and real aircraft models in existing technologies. The device includes: The fault mode setting module is used to set the fault modes of the ailerons. It sets the deflection rate ratio and control surface target deflection for each aileron fault mode. The aileron fault modes include jamming, half-speed deflection, return to center, sharp deflection, damped and non-command oscillation. The control surface deflection output module is used to establish a control surface actuator model. It inputs the control surface hinge torque, the current control surface deflection, the target control surface deflection, and the deflection rate ratio into the control surface actuator model, and outputs the control surface deflection of each aileron through the control surface actuator model. The six-component aerodynamic parameter estimation module is used to obtain the six-component aerodynamic parameters of a single aileron. If the parameters cannot be obtained, the six-component aerodynamic parameters of a single aileron are estimated by using the three-component aerodynamic parameters of the aileron as a whole. The asymmetric correction module is used to perform propeller slipstream asymmetric correction on the six-component aerodynamic parameters to generate asymmetric six-component aerodynamic parameters. The state simulation module is used to calculate the state parameters of the aircraft at the next moment by using the aircraft's six-degree-of-freedom simulation calculation model based on the control surface deflection of each aileron and the asymmetric six-component aerodynamic parameters of the aircraft at the previous moment.
[0007] This invention also provides a computer device, including a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, it implements the above-mentioned simulation method for aileron failure of any medium-sized propeller aircraft, thereby solving the technical problems of coarse simulation models and poor consistency between theoretical simulations and real aircraft models in the prior art for aileron failure.
[0008] This invention also provides a computer-readable storage medium storing a computer program that performs the simulation method for aileron failure of any of the above-described medium-sized propeller aircraft, in order to solve the technical problems in the prior art where the simulation model for aileron failure is coarse and the consistency between theoretical simulation and real aircraft model is poor.
[0009] Compared with the prior art, the beneficial effects that at least one technical solution adopted in the embodiments of this specification can achieve include at least: This solves the problems of existing aileron failure models being coarse and poor consistency between theoretical simulations and real aircraft models. Attached Figure Description
[0010] To more clearly illustrate the technical solutions of the embodiments of this application, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.
[0011] Figure 1 This is a flowchart of a simulation method for aileron failure of a medium-sized propeller aircraft provided in an embodiment of the present invention; Figure 2 This is a flowchart of a simulation method for implementing the above-described aileron failure of a medium-sized propeller aircraft, provided by an embodiment of the present invention. Figure 3 This is a structural block diagram of a computer device provided in an embodiment of the present invention; Figure 4 This is a structural block diagram of a simulation device for aileron failure of a medium-sized propeller aircraft provided in an embodiment of the present invention. Detailed Implementation
[0012] The embodiments of this application will now be described in detail with reference to the accompanying drawings.
[0013] The following specific examples illustrate the implementation of this application. Those skilled in the art can easily understand other advantages and effects of this application from the content disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of this application, and not all of them. This application can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed based on different viewpoints and applications without departing from the spirit of this application. It should be noted that, in the absence of conflict, the following embodiments and features in the embodiments can be combined with each other. Based on the embodiments in this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.
[0014] In this embodiment of the invention, a simulation method for aileron failure in a medium-sized propeller aircraft is provided, such as... Figure 1 As shown, the method includes: Step S101: Set the aileron failure modes, and set the deflection rate ratio and control surface target deflection for each aileron failure mode. The aileron failure modes include jamming, half-speed deflection, return to center, sharp deflection, damped and non-command oscillation. Step S102: Establish a control surface actuator model, input the control surface hinge torque, the current control surface deflection, the target control surface deflection, and the deflection rate ratio into the control surface actuator model, and output the control surface deflection of each aileron through the control surface actuator model; Step S103: Obtain the six-component aerodynamic parameters of a single aileron. If the parameters cannot be obtained, estimate the six-component aerodynamic parameters of a single aileron using the three-component aerodynamic parameters of the aileron as a whole. Step S104: Perform propeller slipstream asymmetric correction on the six-component aerodynamic parameters to generate asymmetric six-component aerodynamic parameters; Step S105: Based on the control surface deflection of each aileron and the asymmetric six-component aerodynamic parameters, the state parameters of the aircraft at the next moment are obtained by simulating the state parameters of the aircraft at the previous moment using the six-degree-of-freedom simulation calculation model.
[0015] In practice, the following steps are used to set the deflection rate ratio and control surface target deflection for each aileron failure mode: The system acquires the target command output by the aircraft control law and sets the sharp yaw angle, damped target yaw angle, and non-command oscillation yaw angle. When the aileron's fault mode is "stuck," the yaw rate ratio is set to 0, and the control surface target deflection is set to the target command. When the aileron's fault mode is "half-speed yaw," the yaw rate ratio is set to 0.5, and the control surface target deflection is set to the target command. When the aileron's fault mode is "return to center," the yaw rate ratio is set to 1, and the control surface target deflection is set to 0. When the aileron's fault mode is "sharp yaw," the yaw rate ratio is set to 1, and the control surface target deflection is set to the sharp yaw angle. When the aileron's fault mode is "damped," the yaw rate ratio is set to 1, and the control surface target deflection is set to the damped target yaw angle. When the aileron's fault mode is "non-command oscillation," the yaw rate ratio is set to 1, and the control surface target deflection is set to the non-command oscillation yaw angle.
[0016] In practice, the following steps are used to estimate the six-component aerodynamic parameters of a single aileron from the three-component aerodynamic parameters of the aileron as a whole: Obtain the overall side force coefficient of the aileron from the aircraft aerodynamic system. Rolling torque coefficient and yaw moment coefficient Set the side force coefficient of the left aileron. ,in, For the left aileron deflection angle, Set the right aileron deflection angle; set the left aileron rolling moment coefficient. Set the yaw moment coefficient of the left aileron. Set the lift coefficient of the left aileron. =0; Sets the drag coefficient of the left aileron. Set the pitch moment coefficient of the left aileron. Set the side force coefficient of the right aileron. Set the rolling moment coefficient of the right aileron. Set the yaw moment coefficient of the right aileron. Set the lift coefficient of the right aileron. Set the drag coefficient for the right aileron. Set the pitch moment coefficient of the right aileron. .
[0017] In specific implementation, the following steps are used to calculate the aircraft's state parameters for the next moment using a six-degree-of-freedom simulation model based on the control surface deflection of each aileron and the asymmetric six-component aerodynamic parameters from the previous moment: By inputting the deflection of each aileron's control surface and the six-component aerodynamic parameters, the net external force and net external moment acting on the aircraft are calculated, where, , The net external force acting on the aircraft. The net external torque acting on the aircraft. , , , and These are the acceleration vector, velocity vector, angular velocity vector, angular momentum vector, and moment of inertia matrix, respectively. As an inertial frame of reference, For the aircraft body shaft system, m For mass; the state parameters of the aircraft at the previous moment, the resultant external force and the resultant external torque are input into the six-degree-of-freedom solution component to obtain the state parameters of the aircraft at the next moment. The state parameters include the three-axis angular rate, Euler angle, velocity, angle of attack, sideslip, flight altitude, engine throttle lever and other control surface deflection angles.
[0018] In specific implementation, the following steps are used to input the control surface hinge torque, the current control surface deflection, the target control surface deflection, and the deflection rate ratio into the control surface actuator model, and the control surface actuator model outputs the control surface deflection of each aileron: Perform the following operations on a single control surface: calculate the actuator load based on the control surface hinge torque and actuator lever arm; calculate the theoretical speed of the actuator output shaft at the current moment based on the interpolation relationship between the actuator load and the actuator output shaft speed; calculate the angle difference between the target deflection of the control surface and the current deflection of the control surface, determine the deflection direction of the control surface based on the angle difference, multiply the deflection direction by the actuator output shaft speed to obtain the theoretical speed of the output shaft; multiply the theoretical speed of the output shaft by the deflection speed proportionally. The actual deflection speed of the control surface actuator under fault mode is calculated; the actual deflection speed is substituted into the integrator to calculate the control surface actuator stroke at the next moment, wherein the upper and lower limits of the integrator are set as the upper and lower limits of the control surface actuator stroke; the control surface deflection angle at the next moment is calculated by back-calculating the control surface actuator stroke and the corresponding relationship; a tolerance threshold is determined, and when the absolute value of the deflection angle difference is less than the tolerance threshold, the control surface deflection at the next moment is used as the control surface deflection of each aileron, and the control surface of the aileron stops deflecting.
[0019] In practice, the non-command oscillation parameter correction is achieved through the following steps: When the fault mode of the aileron is the uncommanded oscillation, dynamic aerodynamic correction can also be performed on the uncommanded oscillation deflection angle, including: obtaining the current airspeed and flight altitude of the aircraft; querying a preset oscillation frequency-amplitude mapping table based on the current airspeed and flight altitude, and dynamically determining the current uncommanded oscillation frequency and oscillation amplitude; and generating a dynamically changing uncommanded oscillation deflection angle based on the uncommanded oscillation frequency and oscillation amplitude.
[0020] In specific implementation, the propeller slipstream asymmetric correction of the six-component aerodynamic parameters is achieved through the following steps to generate asymmetric six-component aerodynamic parameters: After obtaining the six-component aerodynamic parameters of a single aileron, the six-component aerodynamic parameters are subjected to propeller slipstream asymmetric correction, including: Obtain the engine throttle state and propeller rotation direction; based on the throttle state and propeller rotation direction, query a preset slipstream influence coefficient mapping table to obtain the slipstream influence coefficients of the left and right ailerons; multiply the six-component aerodynamic parameters of the left aileron by the slipstream influence coefficient of the left aileron to generate the asymmetric six-component aerodynamic parameters of the left aileron; multiply the six-component aerodynamic parameters of the right aileron by the slipstream influence coefficient of the right aileron to generate the asymmetric six-component aerodynamic parameters of the right aileron.
[0021] In one embodiment of the present invention, such as Figure 2 As shown, the simulation of aileron failure includes: Step 1: Establish a six-degree-of-freedom simulation model of the aircraft.
[0022] Based on the six-degree-of-freedom equilibrium equations of the aircraft, and using control surface deflection and other necessary aircraft parameters as inputs, a preliminary dynamic simulation calculation model of the aircraft is established.
[0023] The model's inputs are the deflection angle of each aileron, the aircraft's state parameters at the previous moment (such as the three-axis angular rate, Euler angle, speed, angle of attack, sideslip, flight altitude, etc.), and other necessary inputs (such as the deflection angles of the engine throttle lever and other control surfaces), and the output is the aircraft's state parameters at the next moment.
[0024] The equations for calculating the resultant external force and resultant moment are as follows: .
[0025] in, and The net external forces and torques acting on the aircraft. , , , and These are the acceleration vector, velocity vector, angular velocity vector, angular momentum vector, and moment of inertia matrix, respectively. As an inertial frame of reference, For the aircraft body shaft system, m For quality.
[0026] Step 2: Refine the aerodynamic parameters of the aileron.
[0027] Replace the longitudinal three-component aerodynamic parameters of all ailerons with the six-component aerodynamic parameters of each aileron. If no readily available six-component data for a single aileron is available, the six-component aerodynamic parameters are given by an approximation method.
[0028] The side force coefficient, roll moment coefficient, and yaw moment coefficient of a single aileron are set by the ratio of the differential deflection angle of the left and right ailerons. If there are lift coefficients, drag coefficients, and pitch moment coefficients of a single aileron, the six-component aerodynamic data of a single aileron are constructed by comprehensively considering them. If there are no lift coefficients, drag coefficients, and pitch moment coefficients, the three-component data are set to 0.
[0029] Side force coefficient of left aileron ,in, For the left aileron deflection angle, The deflection angle of the right aileron; Roll moment coefficient of the left aileron ; Yaw moment coefficient of the left aileron ; Lift coefficient of the left aileron =0; drag coefficient of left aileron ; pitching moment coefficient of the left aileron ; Side force coefficient of the right aileron ; Roll moment coefficient of the right aileron ; Yaw moment coefficient of the right aileron ; Lift coefficient of the right aileron ; drag coefficient of the right aileron ; Pitching moment coefficient of the right aileron .
[0030] Under normal circumstances, and All are fixed values. For example, when the left and right ailerons are deflected at the same angle, it is 0.5. When the left aileron is deflected upward by 1° and the right aileron is deflected downward by 0.5°, it is 2 / 3.
[0031] Step 3: Plan the aileron failure modes.
[0032] Different aileron faults can be set by "setting different fault modes and changing the deflection rate of the actuator and the target deflection angle of the control surface".
[0033] a) Fault numbers are: 0 (normal), 1 (stuck), 2 (half speed deflection), 3 (return to center), 4 (sudden deflection), 5 (damped), 6 (uncommanded oscillation); b) "Rate" represents the deflection rate ratio, which is 0 when "stuck", 0.5 when "half speed deflection", and 1 in other states; c) "Goal" represents the target deflection of the control surface. It is 0 when "returning to center", "deflex" when "sharp", "damping" when "damped", and "Oscillation" when "non-command oscillation". The other states are the target command "KZL" output by the aircraft control law.
[0034] During non-command oscillation, the aircraft's current airspeed and flight altitude are obtained; based on the airspeed and altitude, a preset oscillation frequency-amplitude mapping table is consulted to dynamically determine the oscillation frequency and amplitude; and a dynamically changing non-command oscillation deflection angle is generated based on the above parameters.
[0035] This invention enhances the realism and dynamic adaptability of non-command oscillation fault simulation. In traditional simulations, non-command oscillations are often simplified to periodic signals with fixed frequency and amplitude, failing to reflect the dynamic behavior of oscillation characteristics as airspeed and altitude change during real flight. This invention achieves automatic adjustment of oscillation parameters according to flight conditions through a dynamic mapping table, significantly improving the consistency between the simulation model and real fault behavior. It also enhances the coverage and predictive ability of fault simulation scenarios. By introducing airspeed and altitude as decision variables for oscillation parameters, this method can simulate non-command oscillation characteristics under different flight phases (such as takeoff, cruise, and approach), providing a more comprehensive theoretical basis for flight manual writing and emergency response procedure development. This dynamic correction method does not require complex real-time aeroelastic calculations; it can be achieved simply through table lookup and interpolation, balancing simulation accuracy and computational efficiency, making it suitable for engineering real-time simulation systems.
[0036] Step 4: Integrate the aileron failure modes, establish the control surface actuator model, and output the control surface deflection of each aileron.
[0037] Based on the correspondences of "actuator hinge torque - actuator load", "actuator load - actuator output shaft rate", "control surface target deflection angle - control surface current deflection angle", and "actuator stroke - control surface deflection", and combined with the deflection rate and control surface target deflection angle of the aileron under different fault modes, an aileron actuator model is established.
[0038] It also outputs the real-time control surface deflection angle for each aileron.
[0039] The inputs to the control surface actuator include the control surface hinge torque (Mh), the current control surface deflection (Now), the target control surface deflection (Goal), and the deflection rate ratio (Rate); the output is the control surface deflection angle of each aileron at the next moment (Next).
[0040] The process of constructing the actuator model is as follows: a) The load on the actuator is calculated based on the hinge moment Mh of the aileron and the actuator lever arm Arm. b) The theoretical speed of the aileron actuator output shaft at this moment, |SpeedTheory|, is obtained by interpolation based on the relationship between the load on the aileron actuator and the output shaft speed of the aileron actuator. c) Determine the deflection direction of the aileron control surface based on the difference Delta between the target deflection angle Goal and the current deflection angle Now of the aileron control surface, and multiply it by the aileron actuator output shaft speed to obtain the theoretical speed SpeedTheory of the output shaft. d) Multiply the aileron output shaft theoretical speed SpeedTheory by the deflection rate ratio Rate to obtain the actual deflection speed Speed of the aileron actuator in the fault mode; e) Substitute the actual deflection speed Speed into the integrator to obtain the actuator stroke at the next moment. The upper and lower limits of the integrator are the upper and lower limits of the actuator stroke. f) Calculate the control surface deflection angle Next at the next moment by using the correspondence between the aileron control surface actuator stroke and the control surface deflection; g) Determine the convergence criterion: when the absolute value of the difference between the target value and the current value, Delta, is less than a tolerance value, DeadZone, output the target deflection angle Goal on the rudder surface, and the rudder deflection stops.
[0041] Step 5: Perform propeller slipflow asymmetric correction on the six-component aerodynamic parameters to generate asymmetric six-component aerodynamic parameters.
[0042] Obtain the engine throttle status and propeller rotation direction; query the slipstream influence coefficient mapping table to obtain the slipstream influence coefficients of the left and right ailerons respectively; perform asymmetric correction on the aerodynamic parameters of the left and right ailerons to generate asymmetric six-component aerodynamic parameters.
[0043] This invention accurately simulates the asymmetric impact of propeller slipflow on aileron effectiveness. Addressing the unique slipflow effect of propeller-driven aircraft, it introduces differentiated aerodynamic corrections for the left and right ailerons in fault simulation for the first time. This fully considers the asymmetric impact of slipflow on the aerodynamic environment of the two ailerons, overcoming the limitations of traditional simulations that either ignore slipflow or only treat it symmetrically. It improves simulation accuracy under asymmetric conditions such as single-engine failures, especially when a single engine fails or engine power is asymmetrical, where the propeller slipflow significantly affects the aileron's control effectiveness. By correlating throttle state and rotation direction, this method achieves a dynamic response to the asymmetric slipflow effect, greatly enhancing the predictive reliability of fault simulation under critical conditions. It also enhances the consistency between the simulation model and the aerodynamic characteristics of real aircraft. By introducing a slipflow influence coefficient mapping table, the modeling of complex flow fields is transformed into engineering-implementable parametric corrections. Without significantly increasing the computational burden, it effectively improves the physical fidelity of the aerodynamic model, providing a more reliable simulation platform for flight safety assessment.
[0044] Step 6: Connect to simulation calculations and perform aileron fault simulation.
[0045] The real-time control surface deflection angle of each aileron is output and input into the six-degree-of-freedom simulation model of the aircraft to perform theoretical simulation of aileron failure.
[0046] In this embodiment, a computer device is provided, such as... Figure 3 As shown, it includes a memory 301, a processor 302, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, it implements the above-mentioned simulation method for any medium-sized propeller aircraft aileron failure.
[0047] Specifically, the computer device can be a computer terminal, a server, or a similar computing device.
[0048] In this embodiment, a computer-readable storage medium is provided, which stores a computer program that performs the simulation method for any of the above-described medium-sized propeller aircraft aileron failures.
[0049] Specifically, computer-readable storage media include both permanent and non-permanent, removable and non-removable media, which can store information using any method or technology. Information can be computer-readable instructions, data structures, modules of programs, or other data. Examples of computer-readable storage media include, but are not limited to, phase-change memory (PRAM), static random access memory (SRAM), dynamic random access memory (DRAM), other types of random access memory (RAM), read-only memory (ROM), electrically erasable programmable read-only memory (EEPROM), flash memory or other memory technologies, CD-ROM, digital versatile optical disc (DVD) or other optical storage, magnetic tape, magnetic disk storage or other magnetic storage devices, or any other non-transferable medium that can be used to store information accessible by a computing device. As defined herein, computer-readable storage media do not include transient media, such as modulated data signals and carrier waves.
[0050] Based on the same inventive concept, this invention also provides a simulation device for aileron failure in a medium-sized propeller aircraft, as described in the following embodiments. Since the principle of the simulation device for aileron failure in a medium-sized propeller aircraft is similar to the simulation method for aileron failure in a medium-sized propeller aircraft, the implementation of the simulation device can refer to the implementation of the simulation method for aileron failure in a medium-sized propeller aircraft, and repeated details will not be elaborated further. As used below, the terms "unit" or "module" can refer to a combination of software and / or hardware that performs a predetermined function. Although the device described in the following embodiments is preferably implemented in software, hardware implementation, or a combination of software and hardware, is also possible and contemplated.
[0051] Figure 4 This is a structural block diagram of a simulation device for aileron failure in a medium-sized propeller aircraft according to an embodiment of the present invention, such as... Figure 4 As shown, it includes: a fault mode setting module 401, a control surface deflection output module 402, a six-component aerodynamic parameter estimation module 403, an asymmetric correction module 404, and a state simulation module 405. The structure is described below.
[0052] The fault mode setting module 401 is used to set the fault modes of the aileron. It sets the deflection rate ratio and control surface target deflection for each aileron fault mode. The aileron fault modes include jamming, half-speed deflection, return to center, sharp deflection, damped and non-command oscillation. The control surface deflection output module 402 is used to establish a control surface actuator model, input the control surface hinge torque, the current control surface deflection, the target control surface deflection, and the deflection rate ratio to the control surface actuator model, and output the control surface deflection of each aileron through the control surface actuator model. The six-component aerodynamic parameter estimation module 403 is used to obtain the six-component aerodynamic parameters of a single aileron. If the parameters cannot be obtained, the six-component aerodynamic parameters of a single aileron are estimated by using the three-component aerodynamic parameters of the aileron as a whole. The asymmetric correction module 404 is used to perform propeller slipstream asymmetric correction on the six-component aerodynamic parameters to generate asymmetric six-component aerodynamic parameters. The state simulation module 405 is used to calculate the state parameters of the aircraft at the next moment by using the aircraft's six-degree-of-freedom simulation calculation model based on the control surface deflection of each aileron and the asymmetric six-component aerodynamic parameters of the aircraft at the previous moment.
[0053] In one embodiment, the fault mode setting module includes: The data setting unit is used to acquire the target command output by the aircraft control law and set the sharp yaw angle, damped target yaw angle, and non-command oscillation yaw angle. The jamming setting unit is used to set the deflection rate ratio to 0 and the control surface target deflection to the target command when the aileron's fault mode is jamming. A deflection half-speed setting unit is used to set the deflection rate ratio to 0.5 and the control surface target deflection to the target command when the aileron's fault mode is the deflection half-speed. The centering setting unit is used to set the deflection rate ratio to 1 and the control surface target deflection to 0 when the aileron's failure mode is the centering. A yaw setting unit is used to set the deflection rate ratio to 1 and the control surface target deflection to the yaw angle when the aileron's fault mode is the yaw. The damping setting unit is used to set the deflection rate ratio to 1 and the control surface target deflection to the damping target deflection angle when the aileron's failure mode is the damping. The non-command oscillation setting unit is used to set the deflection rate ratio to 1 and the control surface target deflection to the non-command oscillation deflection angle when the aileron's fault mode is the non-command oscillation.
[0054] In one embodiment, the control surface deflection output module includes: The loop unit is used to perform the following operations on a single control surface: The actuator load calculation unit is used to calculate the actuator load based on the rudder hinge torque and the actuator lever arm. The theoretical rate calculation unit is used to calculate the theoretical rate of the actuator output shaft at the current moment by interpolation based on the correspondence between the actuator load and the actuator output shaft rate. The theoretical speed calculation unit is used to calculate the angle difference between the target deflection of the control surface and the current deflection of the control surface, determine the deflection direction of the control surface based on the angle difference, and multiply the deflection direction by the actuator output shaft speed to obtain the theoretical speed of the output shaft. The actual deflection speed calculation unit is used to multiply the theoretical speed of the output shaft by the deflection rate ratio to calculate the actual deflection speed of the control surface actuator under fault mode. The actuator stroke calculation unit is used to substitute the actual deflection speed into the integrator to calculate the stroke of the control surface actuator at the next moment, wherein the upper and lower limits of the integrator are set as the upper and lower limits of the stroke of the control surface actuator. The control surface deflection angle back calculation unit is used to back calculate the control surface deflection angle at the next moment by using the control surface actuator stroke and the corresponding relationship. The control surface deflection calculation unit is used to determine the tolerance threshold. When the absolute value of the deflection angle difference is less than the tolerance threshold, the control surface deflection at the next moment is used as the control surface deflection of each aileron, and the control surface of the aileron stops deflecting.
[0055] In one embodiment, the six-component aerodynamic parameter estimation module includes: The data acquisition unit is used to obtain the overall side force coefficient of the aileron from the aircraft aerodynamic system. Rolling torque coefficient and yaw moment coefficient ; The left aileron side force coefficient unit is used to set the side force coefficient of the left aileron. ,in, For the left aileron deflection angle, The deflection angle of the right aileron; Set the rolling moment coefficient unit for the left aileron, and set the rolling moment coefficient for the left aileron. ; The unit for setting the yaw moment coefficient of the left aileron is used to set the yaw moment coefficient of the left aileron. ; The lift coefficient unit for the left aileron is used to set the lift coefficient of the left aileron. =0; The drag coefficient unit for the left aileron is used to set the drag coefficient of the left aileron. ; The unit for setting the pitch moment coefficient of the left aileron is used to set the pitch moment coefficient of the left aileron. ; The right aileron side force coefficient unit is used to set the side force coefficient of the right aileron. ; The unit for setting the rolling moment coefficient of the right aileron is used to set the rolling moment coefficient of the right aileron. ; The unit for setting the yaw moment coefficient of the right aileron is used to set the yaw moment coefficient of the right aileron. ; The lift coefficient unit for the right aileron is used to set the lift coefficient of the right aileron. ; The drag coefficient setting unit for the right aileron is used to set the drag coefficient of the right aileron. ; The unit for setting the pitch moment coefficient of the right aileron is used to set the pitch moment coefficient of the right aileron. .
[0056] In one embodiment, the asymmetric correction module includes: The rotation direction acquisition unit is used to acquire the engine throttle status and the propeller rotation direction; The slipstream influence coefficient acquisition unit is used to query a preset slipstream influence coefficient mapping table based on the throttle state and the rotation direction of the propeller to obtain the slipstream influence coefficient of the left aileron and the slipstream influence coefficient of the right aileron. The left aileron correction unit is used to multiply the six-component aerodynamic parameters of the left aileron with the slipstream influence coefficient of the left aileron to generate the asymmetric six-component aerodynamic parameters of the left aileron. The right aileron correction unit is used to multiply the six-component aerodynamic parameters of the right aileron with the slipstream influence coefficient of the right aileron to generate the asymmetric six-component aerodynamic parameters of the right aileron.
[0057] In one embodiment, the state simulation module includes: The resultant external moment calculation unit is used to input the control surface deflection and the six-component aerodynamic parameters of each aileron, and calculate the resultant external force and resultant external moment acting on the aircraft. , The net external force acting on the aircraft. The net external torque acting on the aircraft. , , , and These are the acceleration vector, velocity vector, angular velocity vector, angular momentum vector, and moment of inertia matrix, respectively. As an inertial frame of reference, For the aircraft body shaft system, m For quality; The state calculation unit is used to input the state parameters of the aircraft at the previous moment, the resultant external force, and the resultant external torque into the six-degree-of-freedom calculation component to obtain the state parameters of the aircraft at the next moment. The state parameters include the three-axis angular rate, Euler angle, velocity, angle of attack, sideslip, flight altitude, engine throttle lever, and the deflection angle of other control surfaces.
[0058] In one embodiment, the above-described apparatus further includes a non-command oscillation parameter correction module.
[0059] In one embodiment, the non-command oscillation parameter correction module includes: The data acquisition unit is used to acquire the aircraft's current airspeed and flight altitude; The oscillation unit is used to dynamically determine the current non-command oscillation frequency and oscillation amplitude by querying a preset oscillation frequency-amplitude mapping table based on the current airspeed and flight altitude. The dynamic correction unit is used to generate a dynamically changing non-command oscillation angle based on the non-command oscillation frequency and oscillation amplitude.
[0060] The embodiments of the present invention achieve the following technical effects: This invention provides a simulation calculation method for aileron failure in a medium-sized propeller aircraft, applied to the simulation calculation and analysis after an aileron failure. Compared to existing mathematical models, this invention uses six-component aerodynamic data for each aileron to replace the overall three-component aerodynamic data of the aileron, simulating the impact of each aileron on the aircraft's longitudinal and transverse motions. It adds a control surface actuator model and integrates control surface failure modes, optimizing the control logic and achieving high module integration, making the aileron failure simulation closer to the actual flight control system response and increasing the consistency between the simulation model and the real aircraft. A dynamic mapping table is used to automatically adjust oscillation parameters according to flight conditions, significantly improving the consistency between the simulation model and the actual failure behavior. Differential aerodynamic corrections for the left and right ailerons are introduced into the failure simulation, fully considering the asymmetric impact of the slipstream on the aerodynamic environment of both ailerons, overcoming the limitations of traditional simulations that ignore the slipstream or only perform symmetrical treatment.
[0061] Obviously, those skilled in the art should understand that the modules or steps of the above-described embodiments of the present invention can be implemented using general-purpose computing devices. They can be centralized on a single computing device or distributed across a network of multiple computing devices. Optionally, they can be implemented using computer-executable program code, thereby storing them in a storage device for execution by a computing device. In some cases, the steps shown or described can be performed in a different order than those presented here, or they can be fabricated as separate integrated circuit modules, or multiple modules or steps can be fabricated as a single integrated circuit module. Thus, the embodiments of the present invention are not limited to any particular hardware and software combination.
[0062] The above description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, various modifications and variations can be made to the embodiments of the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.
Claims
1. A simulation method for aileron failure in a medium-sized propeller aircraft, characterized in that, include: Set the failure modes of the aileron, and set the deflection rate ratio and control surface target deflection for each failure mode of the aileron. The failure modes of the aileron include jamming, half-speed deflection, return to center, sharp deflection, damped and non-command oscillation. Establish a control surface actuator model, input the control surface hinge torque, the current control surface deflection, the target control surface deflection, and the deflection rate ratio into the control surface actuator model, and output the control surface deflection of each aileron through the control surface actuator model; Obtain the six-component aerodynamic parameters of a single aileron. If these parameters cannot be obtained, estimate the six-component aerodynamic parameters of a single aileron using the three-component aerodynamic parameters of the aileron as a whole. The six-component aerodynamic parameters are modified by propeller slipflow asymmetry to generate asymmetric six-component aerodynamic parameters; Based on the control surface deflection of each aileron and the asymmetric six-component aerodynamic parameters, the state parameters of the aircraft at the next moment are obtained by simulating the state parameters of the aircraft at the previous moment using the six-degree-of-freedom simulation model.
2. The simulation method for aileron failure in a medium-sized propeller aircraft as described in claim 1, characterized in that, Set the deflection rate ratio and control surface target deflection for each aileron failure mode, including: Obtain the target command output by the aircraft control law, and set the sharp yaw angle, damped target yaw angle, and non-command oscillation yaw angle; When the aileron's fault mode is the jamming, the deflection rate ratio is set to 0, and the control surface target deflection is set to the target command. When the aileron's failure mode is the half-speed deflection, the deflection rate ratio is set to 0.5, and the control surface target deflection is set to the target command. When the aileron's failure mode is the centering out, the deflection rate ratio is set to 1, and the control surface target deflection is set to 0. When the aileron's fault mode is the sharp deflection, the deflection rate ratio is set to 1, and the control surface target deflection is set to the sharp deflection angle. When the aileron's failure mode is the damping, the deflection rate ratio is set to 1, and the control surface target deflection is set to the damping target deflection angle. When the aileron's fault mode is the non-command oscillation, the deflection rate ratio is set to 1, and the control surface target deflection is set to the non-command oscillation deflection angle.
3. The simulation method for aileron failure in a medium-sized propeller aircraft as described in claim 1, characterized in that, The six-component aerodynamic parameters of a single aileron are estimated using the three-component aerodynamic parameters of the aileron as a whole, including: Obtain the overall side force coefficient of the aileron from the aircraft aerodynamic system. Rolling torque coefficient and yaw moment coefficient ; Set the side force coefficient of the left aileron ,in, For the left aileron deflection angle, The deflection angle of the right aileron; Set the rolling moment coefficient of the left aileron ; Set the yaw moment coefficient of the left aileron ; Set the lift coefficient of the left aileron =0; Set the drag coefficient of the left aileron ; Set the pitch moment coefficient of the left aileron. ; Set the side force coefficient of the right aileron ; Set the rolling moment coefficient of the right aileron. ; Set the yaw moment coefficient of the right aileron. ; Set the lift coefficient of the right aileron ; Set the drag coefficient of the right aileron ; Set the pitch moment coefficient of the right aileron. .
4. The simulation method for aileron failure in a medium-sized propeller aircraft as described in claim 1, characterized in that, Based on the control surface deflection of each aileron and the asymmetric six-component aerodynamic parameters, the aircraft's state parameters for the next moment are calculated using a six-degree-of-freedom simulation model through simulation of the aircraft's state parameters at the previous moment, including: By inputting the deflection of each aileron's control surface and the six-component aerodynamic parameters, the net external force and net external moment acting on the aircraft are calculated, where, , The net external force acting on the aircraft. The net external torque acting on the aircraft. , , , and These are the acceleration vector, velocity vector, angular velocity vector, angular momentum vector, and moment of inertia matrix, respectively. As an inertial frame of reference, For the aircraft body shaft system, m For quality; The state parameters of the aircraft at the previous moment, the resultant external force, and the resultant external torque are input into the six-degree-of-freedom solution component to obtain the state parameters of the aircraft at the next moment. The state parameters include the three-axis angular rate, Euler angle, velocity, angle of attack, sideslip, flight altitude, engine throttle lever, and the deflection angle of other control surfaces.
5. The simulation method for aileron failure in a medium-sized propeller aircraft as described in claim 1, characterized in that, The control surface hinge moment, current control surface deflection, target control surface deflection, and deflection rate ratio are input into the control surface actuator model. The control surface actuator model outputs the control surface deflection of each aileron, including: Perform the following operations on a single control surface: The load on the actuator is calculated based on the hinge torque of the rudder surface and the actuator lever arm. The theoretical speed of the actuator output shaft at the current moment is calculated by interpolation based on the correspondence between the load on the actuator and the speed of the actuator output shaft. Calculate the angle difference between the target deflection of the control surface and the current deflection of the control surface, determine the deflection direction of the control surface based on the angle difference, and multiply the deflection direction by the actuator output shaft rate to obtain the theoretical speed of the output shaft. The actual deflection speed of the control surface actuator under fault mode is calculated by multiplying the theoretical speed of the output shaft by the ratio of the deflection rate. The actual deflection speed is substituted into the integrator to calculate the stroke of the control surface actuator at the next moment, wherein the upper and lower limits of the integrator are set as the upper and lower limits of the stroke of the control surface actuator. The deflection angle of the control surface at the next moment can be calculated by inversely using the corresponding relationship between the travel of the control surface actuator and the control surface. A tolerance threshold is determined. When the absolute value of the deflection angle difference is less than the tolerance threshold, the control surface deflection at the next moment is taken as the control surface deflection of each aileron, and the control surface of the aileron stops deflecting.
6. The simulation method for aileron failure of a medium-sized propeller aircraft as described in any one of claims 1 to 5, characterized in that, Also includes: When the aileron's failure mode is the uncommanded oscillation, dynamic aerodynamic correction can also be performed on the uncommanded oscillation deflection angle, including: Obtain the aircraft's current airspeed and altitude; Based on the current airspeed and flight altitude, the preset oscillation frequency-amplitude mapping table is queried to dynamically determine the current non-command oscillation frequency and oscillation amplitude; Based on the non-command oscillation frequency and oscillation amplitude, a dynamically changing non-command oscillation deflection angle is generated.
7. The simulation method for aileron failure of a medium-sized propeller aircraft as described in any one of claims 1 to 5, characterized in that, The six-component aerodynamic parameters are subjected to propeller slipflow asymmetric correction to generate asymmetric six-component aerodynamic parameters, including: After obtaining the six-component aerodynamic parameters of a single aileron, the six-component aerodynamic parameters are subjected to propeller slipstream asymmetric correction, including: Obtain the engine throttle status and propeller rotation direction; Based on the throttle state and the rotation direction of the propeller, a preset slipstream influence coefficient mapping table is consulted to obtain the slipstream influence coefficients of the left aileron and the right aileron. Multiply the six-component aerodynamic parameters of the left aileron by the slipstream influence coefficient of the left aileron to generate the asymmetric six-component aerodynamic parameters of the left aileron; The asymmetric six-component aerodynamic parameters of the right aileron are generated by multiplying the six-component aerodynamic parameters of the right aileron by the slipstream influence coefficient of the right aileron.
8. A simulation device for aileron failure in a medium-sized propeller aircraft, characterized in that, include: The fault mode setting module is used to set the fault modes of the ailerons. It sets the deflection rate ratio and control surface target deflection for each aileron fault mode. The aileron fault modes include jamming, half-speed deflection, return to center, sharp deflection, damped and non-command oscillation. The control surface deflection output module is used to establish a control surface actuator model. It inputs the control surface hinge torque, the current control surface deflection, the target control surface deflection, and the deflection rate ratio into the control surface actuator model, and outputs the control surface deflection of each aileron through the control surface actuator model. The six-component aerodynamic parameter estimation module is used to obtain the six-component aerodynamic parameters of a single aileron. If the parameters cannot be obtained, the six-component aerodynamic parameters of a single aileron are estimated by using the three-component aerodynamic parameters of the aileron as a whole. The asymmetric correction module is used to perform propeller slipstream asymmetric correction on the six-component aerodynamic parameters to generate asymmetric six-component aerodynamic parameters. The state simulation module is used to calculate the state parameters of the aircraft at the next moment by using the aircraft's six-degree-of-freedom simulation calculation model based on the control surface deflection of each aileron and the asymmetric six-component aerodynamic parameters of the aircraft at the previous moment.
9. A computer device, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that, When the processor executes the computer program, it implements a simulation method for aileron failure of a medium-sized propeller aircraft as described in any one of claims 1 to 7.
10. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores a computer program that performs a simulation method for aileron failure of a medium-sized propeller aircraft as described in any one of claims 1 to 7.