Large aircraft actuator mode conversion control law reconstruction method
By analyzing and reconstructing the control law after a mode transition failure of a large aircraft actuator, and adjusting the control commands and gains of the control surfaces, the problems of damaged control surface deflection logic and integrator saturation caused by the mode transition failure were solved, thereby improving flight safety and the tolerance of the control law.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-12-31
- Publication Date
- 2026-04-14
AI Technical Summary
When a large aircraft's actuators fail to switch modes, they are unable to execute normal mode control law commands, resulting in damage to the control surface deflection logic and saturation of the control law integrator, which affects flight quality and safety.
By analyzing the impact of actuator mode conversion failure, the control law reconstruction principle is determined, the control surface control commands and gains are adjusted, the faulty function is disconnected, and the normal control surface is used to replace the faulty control surface to ensure the control surface deflection logic and flight quality.
This technology improves the fault tolerance of control laws without changing the hardware, ensuring flight safety and the security of control law calculations, while reducing modification costs.
Smart Images
Figure CN121857792A_ABST
Abstract
Description
Technical Field
[0001] This application belongs to the field of flight control system technology, and specifically relates to a method for reconstructing the mode conversion control law of a large aircraft actuator. Background Technology
[0002] Large aircraft flight control systems are designed based on high safety principles. Actuator control modes include normal and backup modes, receiving different control commands in each mode. When an actuator experiences a mode transition failure, it switches from receiving normal mode control commands to receiving backup mode control commands. This prevents the faulty actuator from executing normal mode control law commands, thus hindering the normal mode control law function. Furthermore, it disrupts the control surface deflection logic and can cause control law integrator saturation.
[0003] Therefore, when an actuator mode switching failure occurs, the logic and parameters of the control law need to be reconstructed to ensure the control surface deflection logic and flight quality, ensure the safety of control law calculation, and improve the control law fault tolerance. Summary of the Invention
[0004] The purpose of this application is to provide a method for reconstructing the mode conversion control law of a large aircraft actuator, so as to solve or alleviate at least one of the problems in the prior art.
[0005] The technical solution of this application is: a method for reconstructing the mode conversion control law of a large aircraft actuator, comprising:
[0006] Step S1: Analyze the impact of mode transition failures in various actuators of a large aircraft and obtain the analysis results.
[0007] Step 2: Determine the reconstructing principle of the actuator mode conversion control law based on the analysis results, and reconstruct the control law based on the reconstructing principle.
[0008] Preferably, the analysis result is:
[0009] When some elevator actuators experience mode switching failure, the remaining normal elevators will not be effective enough to control the aircraft's pitching or slamming motion, causing the integrator commands to accumulate until saturation. When the stabilizer actuators experience mode switching failure, they will not be able to replace the elevators for trim control after the aircraft has entered steady-state flight.
[0010] When some rudder actuators experience mode switching failures, the coordinated sideslip function will be affected; when one aileron actuator on one side of the wing experiences a mode switching failure, the aileron's deflection logic will be disrupted.
[0011] Preferably, the reconstruction principle includes:
[0012] a) When there are many actuators undergoing mode switching, the corresponding functions cannot be executed. The actuator functions should be disconnected and the pilot should be given a function loss warning.
[0013] b) Adjust the control commands for similar normal control surfaces to ensure the deflection logic as much as possible;
[0014] c) Adjust the control gain of the remaining similar normal control surfaces to maintain the desired control effect;
[0015] d) Adjust the control logic of the control surfaces with the same function so that the normal control surface can replace the faulty control surface, thereby ensuring that the desired function is achieved.
[0016] Preferably, based on reconstruction principle a, the elevator actuator mode transition control law reconstruction method is as follows:
[0017] When half or more of the elevator actuators undergo mode switching, the automatic trim, pitch limit, normal overload limit, and longitudinal functions can no longer achieve normal control. In order to enable the pilot to understand the situation and avoid integrator deep saturation, the longitudinal function is disconnected and the pilot is given a function loss warning.
[0018] Preferably, based on reconstruction principle b, the aileron actuator mode transition control law reconstruction method is as follows:
[0019] After a mode transition occurs in the aileron actuator on one side of the wing, in order to ensure that the deflection of the left and right ailerons on both sides of the wing conforms to the control surface usage logic as much as possible, the normal mode lateral control law is converted into direct chain control, that is, the control surface deflection angle corresponds to the control command, thereby matching the deflection angle of the faulty aileron.
[0020] Preferably, based on reconstruction principle c, the reconstructing method for the rudder actuator mode transition control law is as follows:
[0021] After the rudder actuator undergoes a mode transition, the normal rudder deflection angle is increased by increasing the gain of the coordinated sideslip control branch, thereby ensuring the automatic sideslip reduction effect of the rudder when the aircraft rolls; if there is no normal rudder, no fault reconstruction is performed, and the pilot is given an alarm indicating the loss of coordinated sideslip function.
[0022] Preferably, based on the reconfiguration principle d, the reconfiguration method for the mode transition control law of the trimmed rudder actuator is as follows:
[0023] When the horizontal stabilizer actuator undergoes a mode transition, the control circuit of the horizontal stabilizer, which replaces the elevator, is disconnected, and the elevator performs the trim function.
[0024] The large aircraft actuator mode transition control law reconstruction method disclosed in this application reconstructs the control law logic and parameters after an actuator mode transition failure, thereby ensuring control surface deflection logic and flight quality, guaranteeing the safety of control law calculation, and improving the control law fault tolerance capability. This method does not require any modification to the flight control system hardware. Actuator mode transition reconstruction control laws can be designed according to this method, saving modification costs and improving aircraft safety after actuator mode transition failures in large aircraft. Attached Figure Description
[0025] To more clearly illustrate the technical solutions provided in this application, the accompanying drawings will be briefly described below. Obviously, the drawings described below are merely some embodiments of this application.
[0026] Figure 1 This is a schematic diagram of the control reconfiguration method for the mode conversion of large aircraft actuators according to this application. Detailed Implementation
[0027] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below with reference to the accompanying drawings.
[0028] This application proposes a control law reconstruction method for the mode transition failure of large aircraft actuators. The method reconstructs the logic and parameters of the control law to ensure the control surface deflection logic and flight quality, ensure the safety of control law calculation, and improve the control law fault tolerance.
[0029] like Figure 1 As shown, the large aircraft actuator mode conversion control law reconstruction method provided in this application includes the following process:
[0030] Step S1: Analyze the impact of mode transition failures in various actuators of a large aircraft and obtain the analysis results.
[0031] The flight control system of large aircraft is designed based on high safety principles and is divided into normal mode and backup mode. Actuators receive different control commands in different flight control system modes. These commands are calculated and generated by the normal mode control law and the backup mode control law in the flight control computer. The normal mode control law uses complex control laws to achieve complete flight control system functions, while the backup mode control law uses simpler control laws to achieve aircraft controllability. When an actuator undergoes a mode transition due to a malfunction—that is, when the actuator switches from receiving normal mode control commands to receiving backup mode control commands—the malfunctioning actuator will be unable to execute the normal mode control law commands.
[0032] For large aircraft, the horizontal stabilizer is equipped with elevators (e.g., two or four elevators) that are controlled by the flight control computer to deflect upwards or downwards to control the aircraft's pitching or yaw. The horizontal stabilizer is also equipped with a horizontal stabilizer that is controlled by the flight control computer to deflect upwards or downwards to control the aircraft's pitching or yaw. The vertical stabilizer is equipped with rudders (e.g., one or two rudders) that are controlled by the flight control computer to deflect upwards or downwards to control the aircraft's yaw to the left or right. The left and right wings are equipped with left and right ailerons, respectively. These ailerons are controlled by the flight control computer to perform differential movements, such as deflecting the left aileron upwards and the right aileron downwards, or deflecting the left aileron downwards and the right aileron upwards (from the forward view at the tail), thereby controlling the aircraft's roll to the left or right. Each control surface is driven by actuators.
[0033] The normal modal control law employs a proportional-integral feedback control principle, achieving functions such as automatic trim, pitch angle limiting, and normal overload limiting. The key component is the integrator, which continuously accumulates elevator deflection commands when there is a deviation between the aircraft's response and control commands. In large aircraft control modes, the stabilizer plays a trim role, replacing the elevator's trim angle once the aircraft reaches steady-state flight. The rudder coordinates sideslip, automatically deflecting to eliminate sideslip during aircraft roll. To eliminate the adverse yaw moment generated by aileron deflection, the aileron's upward deflection angle is greater than its downward deflection angle, for example, the upward deflection angle is 1.5 times the downward deflection angle.
[0034] Therefore, based on the above analysis of the functions achieved by the control surfaces, it can be concluded that:
[0035] When some elevator actuators experience mode switching failure, they switch from receiving normal mode control law commands to receiving backup mode control law commands. In order to achieve functions such as automatic trim, pitch angle limiting, and normal overload limiting, the deflection angle of the remaining normal elevators will increase. If there are many faulty elevators, the effectiveness of the remaining normal elevators will be insufficient to achieve the above functions, resulting in a continuous deviation between the aircraft response and control commands, leading to the accumulation of integrator commands until saturation. When the stabilizer actuators experience mode switching failure, they will be unable to replace the elevators for trim control after the aircraft has entered steady-state flight.
[0036] When some rudder actuators experience mode switching failures, the coordinated sideslip function will be affected; when one aileron actuator on one side of the wing experiences a mode switching failure, the deflection logic of the aileron will be disrupted due to the significant difference between the commands of the normal mode control law and the backup mode control law.
[0037] Step S2: Determine the reconstructing principle of the actuator mode conversion control law based on the analysis results, and reconstruct the control law based on the reconstructing principle.
[0038] In this application, the actuator mode transition control law reconfiguration principle includes:
[0039] a) Disconnect functions for safety: When there are many actuators undergoing mode switching, the corresponding functions cannot be executed. These functions should be disconnected and the pilot should be given a function loss warning.
[0040] b) Switching commands to maintain logic: Adjusting similar normal control commands for the control surfaces to ensure deflection logic as much as possible;
[0041] c) Adjust gain to maintain effect: Adjust the control gain of the remaining normal control surfaces of the same type to maintain the desired control effect;
[0042] d) Reconstruct logic to ensure functionality: Adjust the control logic of the control surfaces with the same function so that the normal control surface can replace the faulty control surface, thereby ensuring that the desired function is achieved.
[0043] Based on control law reconstruction principle a, the elevator actuator mode conversion control law reconstruction method in this application is as follows: after half or more of the elevator actuators undergo mode conversion, the longitudinal functions such as automatic trim, pitch angle limit, and normal overload limit can no longer achieve normal control effects. In order to enable the pilot to understand the situation and avoid integrator deep saturation, it is necessary to disconnect these longitudinal functions and give the pilot a function loss alarm prompt.
[0044] Based on control law reconstruction principle b, the control law reconstruction method for aileron actuator mode conversion in this application is as follows: after aileron actuator on one side of the wing undergoes mode conversion, in order to ensure that the deflection of the left and right ailerons on both sides of the wing conforms to the control surface usage logic as much as possible, the normal mode lateral control law should be converted into direct chain control, that is, the control surface deflection angle corresponds to the control command, thereby matching the faulty aileron deflection angle.
[0045] Based on control law reconstruction principle c, the rudder actuator mode conversion control law reconstruction method in this application is as follows: after the rudder actuator undergoes mode conversion, the gain of the coordinated sideslip control branch can be increased, thereby increasing the normal rudder deflection angle and ensuring the automatic sideslip elimination effect of the rudder when the aircraft rolls; if there is no normal rudder, fault reconstruction can be omitted, and the pilot will be given an alarm prompt of loss of coordinated sideslip function.
[0046] Based on the control law reconstruction principle d, the control law reconstruction method for the trim surface actuator mode conversion in this application is as follows: when the horizontal stabilizer actuator undergoes mode conversion, the control loop of the horizontal stabilizer replacing the elevator needs to be disconnected, and the elevator performs the trim function.
[0047] The large aircraft actuator mode transition control law reconstruction method disclosed in this application reconstructs the control law logic and parameters after an actuator mode transition failure, thereby ensuring control surface deflection logic and flight quality, guaranteeing the safety of control law calculation, and improving the control law fault tolerance capability. This method does not require any modification to the flight control system hardware. Actuator mode transition reconstruction control laws can be designed according to this method, saving modification costs and improving aircraft safety after actuator mode transition failures in large aircraft.
[0048] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.
Claims
1. A method for reconstructing the mode conversion control law of a large aircraft actuator, characterized in that, include: Step S1: Analyze the impact of mode transition failures in various actuators of a large aircraft and obtain the analysis results; Step 2: Determine the reconstructing principle of the actuator mode conversion control law based on the analysis results, and reconstruct the control law based on the reconstructing principle.
2. The method for reconstructing the mode conversion control law of a large aircraft actuator as described in claim 1, characterized in that, The analysis results are as follows: When some elevator actuators experience mode switching failure, the remaining normal elevators will not be effective enough to control the aircraft's pitching or slamming motion, causing the integrator commands to accumulate until saturation. When the stabilizer actuators experience mode switching failure, they will not be able to replace the elevators for trim control after the aircraft has entered steady-state flight. When some rudder actuators experience mode switching failures, the coordinated sideslip function will be affected; when one aileron actuator on one side of the wing experiences a mode switching failure, the aileron's deflection logic will be disrupted.
3. The method for reconstructing the mode conversion control law of a large aircraft actuator as described in claim 1, characterized in that, The restructuring principles include: a) When there are many actuators undergoing mode switching, the corresponding functions cannot be executed. The actuator functions should be disconnected and the pilot should be given a function loss warning. b) Adjust the control commands for similar normal control surfaces to ensure the deflection logic as much as possible; c) Adjust the control gain of the remaining similar normal control surfaces to maintain the desired control effect; d) Adjust the control logic of the control surfaces with the same function so that the normal control surface can replace the faulty control surface, thereby ensuring that the desired function is achieved.
4. The method for reconstructing the mode conversion control law of a large aircraft actuator as described in claim 3, characterized in that, Based on reconstruction principle a, the reconstruction method for the elevator actuator mode transition control law is as follows: When half or more of the elevator actuators undergo mode switching, the automatic trim, pitch limit, normal overload limit, and longitudinal functions can no longer achieve normal control. In order to enable the pilot to understand the situation and avoid integrator deep saturation, the longitudinal function is disconnected and the pilot is given a function loss warning.
5. The method for reconstructing the mode conversion control law of a large aircraft actuator as described in claim 3, characterized in that, Based on reconstruction principle b, the reconstruction method for the aileron actuator mode transition control law is as follows: After a mode transition occurs in the aileron actuator on one side of the wing, in order to ensure that the deflection of the left and right ailerons on both sides of the wing conforms to the control surface usage logic as much as possible, the normal mode lateral control law is converted into direct chain control, that is, the control surface deflection angle corresponds to the control command, thereby matching the deflection angle of the faulty aileron.
6. The method for reconstructing the mode conversion control law of a large aircraft actuator as described in claim 3, characterized in that, Based on reconstruction principle c, the reconstructing method for the rudder actuator mode transition control law is as follows: After the rudder actuator undergoes a mode transition, the normal rudder deflection angle is increased by increasing the gain of the coordinated sideslip control branch, thereby ensuring the automatic sideslip reduction effect of the rudder when the aircraft rolls; if there is no normal rudder, no fault reconstruction is performed, and the pilot is given an alarm indicating the loss of coordinated sideslip function.
7. The method for reconstructing the mode conversion control law of a large aircraft actuator as described in claim 3, characterized in that, Based on reconstruction principle d, the reconstructing method for the mode transition control law of the trimmed rudder actuator is as follows: When the horizontal stabilizer actuator undergoes a mode transition, the control circuit of the horizontal stabilizer, which replaces the elevator, is disconnected, and the elevator performs the trim function.