Static instability aircraft actuator torque saturation protection control method and device
By identifying actuator torque saturation in statically unstable aircraft and executing control command attenuation and disconnecting the integrator channel, the aircraft divergence problem caused by actuator torque saturation in statically unstable aircraft is solved, realizing the protective control of the flight control system and avoiding hardware modifications.
Patent Information
- Application Number
- CN202512013616.3
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-12-29
- Publication Date
- 2026-03-24
AI Technical Summary
When the actuator torque is saturated, statically unstable aircraft cannot effectively control the deflection of the control surfaces, resulting in the divergence of the aircraft's motion. Current technology cannot effectively avoid this situation.
By identifying actuator torque saturation, executing control command attenuation, and disconnecting the integrator control channel, a control method and device are designed to protect actuator torque saturation and prevent aircraft divergence.
It effectively avoids aircraft divergence caused by actuator torque saturation in statically unstable aircraft, and does not require modification of flight control system hardware, thus saving costs.
Smart Images
Figure CN121722172A_ABST
Abstract
Description
Technical Field
[0001] This application belongs to the field of aircraft flight control technology, and specifically relates to a method and device for torque saturation protection control of statically unstable aircraft actuators. Background Technology
[0002] When a statically unstable aircraft is manipulated or disturbed, its flight parameters rapidly diverge, requiring stabilization control via the flight control system. This involves maneuvering actuators to drive the deflection of control surfaces, generating stabilizing torque to stabilize the aircraft. If the stabilizing torque generated by the control surface deflection is insufficient to meet the aircraft's stabilization requirements, the aircraft's motion will continue to diverge. Actuators are designed based on maximum output force requirements, derived from calculated hinge torque data obtained through wind tunnel testing. Since calculations and tests cannot perfectly replicate actual flight conditions, actuator output force may sometimes be lower than required under certain flight conditions. This is particularly true for aircraft control surfaces controlled by two sets of actuators simultaneously. If one set of actuators fails, the output force generated by the remaining working actuators will be significantly less than the combined output force of both sets. The aerodynamic forces on the control surfaces generate hinge torque relative to the control surface hinge axis. If the torque generated by the actuator's output force cannot overcome this hinge torque, the actuator will be unable to control the control surface deflection. More seriously, if the torque generated by the actuator's output force is significantly less than the hinge torque, the control surface will deflect in the opposite direction under the influence of aerodynamic forces. These phenomena are called actuator torque saturation. Once the actuator experiences torque saturation, it will be unable to control the control surface to deflect to the desired angle, resulting in the inability to generate the stabilizing torque required by the aircraft, and the aircraft will diverge.
[0003] Therefore, it is necessary to design an actuator torque saturation protection control law for the actuator torque saturation condition of statically unstable aircraft to avoid aircraft divergence caused by actuator torque saturation. Summary of the Invention
[0004] The purpose of this application is to provide a method and apparatus for protecting the torque saturation of statically unstable aircraft actuators, in order to solve or mitigate at least one of the problems in the prior art.
[0005] The technical solution of this application is: a method for torque saturation protection control of actuators in statically unstable aircraft, comprising:
[0006] Identify actuator torque saturation and determine whether actuator torque saturation has occurred;
[0007] When the actuator torque is saturated, an actuator torque saturation protection operation is executed. The actuator torque saturation protection operation includes attenuating the pilot's control commands and disconnecting the integrator control channel, thereby realizing the protection setting of the control law.
[0008] Preferably, the method for determining whether the actuator torque saturation occurs is:
[0009] a) the rudder position and the rudder deflection command have a predetermined deviation M for a time period T;
[0010] b) the rudder position movement speed in the direction of the rudder deflection command is lower than a predetermined speed threshold K, or the rudder position movement direction is opposite to the rudder deflection command direction;
[0011] c) the dynamic pressure signal is greater than a dynamic pressure threshold S;
[0012] When all the above three conditions are met, it is determined that the actuator torque saturation occurs.
[0013] Preferably, the time period T is greater than the maximum response time of the actuator, which is the time from the start to the maximum displacement of the actuator when the maximum displacement command is applied to the actuator.
[0014] Preferably, the deviation M is 5% to 10% of the rudder deflection command.
[0015] Preferably, the speed threshold K is the rudder deflection speed under the maximum hinge torque, which is determined according to the actuator loading test.
[0016] Preferably, the dynamic pressure threshold S is determined by the hinge torque data and the output force performance of the actuator obtained by calculation and wind tunnel test.
[0017] Preferably, when the actuator torque saturation occurs, the process of attenuating the pilot control command is:
[0018] If the actuator torque saturation occurs for the first time, the pilot control command is attenuated by a predetermined proportion, thereby reducing the aircraft movement response and reducing the rudder stability augmentation demand;
[0019] If the actuator torque saturation still occurs, the pilot control command is attenuated by a predetermined proportion on the basis of the first attenuation;
[0020] And so on, until the actuator torque saturation no longer occurs.
[0021] Preferably, the predetermined proportion is 30% to 50%.
[0022] Preferably, when the actuator torque saturation occurs, the method for disconnecting the integrator control channel is:
[0023] By making the integrator channel switch in the "off" state, the integrator control channel is disconnected, thereby avoiding integrator saturation; when it is detected that the actuator torque is not saturated, the integrator control channel is connected by making the integrator channel switch in the "on" state, thereby restoring the normal control architecture.
[0024] In addition, this application also provides an apparatus for implementing the above-described statically unstable aircraft actuator torque saturation protection control method, which includes a forward control channel, a proportional control channel, an integral control channel, an actuator torque saturation identification module, an integrator on / off switch module, and a command attenuation module. The actuator torque saturation identification module is used to identify actuator torque saturation and determine whether actuator torque saturation has occurred. The command attenuation module is used to attenuate the pilot control command when the actuator torque is saturated. The integrator on / off switch module is used to disconnect when the actuator torque is saturated, so as to disconnect the integrator control channel.
[0025] The method and device for protecting the torque saturation of actuators in statically unstable aircraft provided in this application can prevent the aircraft from diverging due to the torque saturation of actuators in statically unstable aircraft. It does not require any modification to the hardware of the flight control system. The actuator torque saturation protection control law can be designed according to this method and added to the flight control law of the flight control system, saving modification costs and realizing the protection of the torque saturation of actuators in statically unstable aircraft. Attached Figure Description
[0026] 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.
[0027] Figure 1 This is a schematic diagram of the torque saturation protection control law for statically unstable aircraft actuators in this application. Detailed Implementation
[0028] 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.
[0029] This application provides a torque saturation protection control method for actuators in statically unstable aircraft to prevent aircraft divergence caused by torque saturation of actuators in statically unstable aircraft.
[0030] like Figure 1 As shown, the torque saturation protection control method for statically unstable aircraft actuators provided in this application includes the following process:
[0031] Step S10: Identify actuator torque saturation and determine whether actuator torque saturation has occurred;
[0032] Identifying actuator torque saturation conditions and designing actuator torque saturation protection mechanisms in the control law can prevent actuator torque saturation from causing aircraft divergence. For example... Figure 1As shown, a typical control law architecture contains a forward control channel, a proportional control channel and an integral control channel. Pilot control commands are passed through these three channels to generate a rudder deflection command, which controls the actuator movement to deflect the rudder.
[0033] In this application, the method for determining actuator torque saturation is:
[0034] a) the rudder position and the rudder deflection command have a deviation M for a time period T;
[0035] b) the rudder position moves at a speed lower than a certain speed threshold K in the direction of the rudder deflection command, or the rudder position moves in the opposite direction of the rudder deflection command;
[0036] c) the dynamic pressure signal is greater than a dynamic pressure threshold S.
[0037] When all the above three conditions are met, it is determined that the actuator is in torque saturation.
[0038] In condition a of this application, the method for selecting the time period T is as follows: the maximum displacement command is applied to the actuator, and the time for the actuator to reach the maximum displacement from the start is recorded, which is the maximum response time of the actuator; within the maximum response time of the actuator, the deviation between the rudder position and the rudder deflection command is normal, but within a sufficient time, if the actuator cannot reach the rudder deflection command, it is considered that the actuator cannot move normally, therefore, the time period T should be greater than the maximum response time of the actuator, for example, T = maximum response time + N, N is a set margin, for example, N can be 0.5s.
[0039] In condition a of this application, the deviation M can generally be selected as 5% to 10% of the rudder deflection command.
[0040] In condition b of this application, if the rudder position moves at a speed lower than a certain threshold in the direction of the rudder deflection command, or the rudder position moves in the opposite direction of the rudder deflection command, it is determined that the actuator is in torque saturation, which cannot overcome the rudder hinge torque, thereby affecting the movement speed and direction of the actuator, resulting in the rudder deflection not meeting the expectations. Therefore, the speed threshold K is determined according to the actuator loading test, and the maximum hinge torque determined by calculation and wind tunnel test is loaded, and the movement speed of the actuator under this load is checked, and the rudder deflection speed obtained by the steady-state speed of the actuator under this load is set as the speed threshold K, i.e. the rudder deflection speed under the maximum hinge torque.
[0041] In this application, the hinge torque is closely related to the dynamic pressure, and only when the dynamic pressure is greater than a certain dynamic pressure threshold S, it is determined that the hinge torque of the rudder is large at this time, therefore, the dynamic pressure threshold S can be determined by the hinge torque data obtained by calculation and wind tunnel test, and the output force performance of the actuator is combined to select the dynamic pressure threshold S when the hinge torque is saturated. For example, when the dynamic pressure is 1000kg / m2 When the rudder surface is deflected at a large angle, the hinge moment is equal to the output force, and the dynamic pressure threshold S is taken as 1000kg / m 2 .
[0042] Step S20, when the actuator torque is saturated, an actuator torque saturation protection operation is performed, which includes attenuating the pilot control command and disconnecting the integrator control channel, so as to achieve the protection setting of the control law.
[0043] Continuing to refer to Figure 1 As shown in the figure, after identifying the actuator torque saturation, the pilot control command is attenuated, and the integrator control channel is disconnected, so as to achieve the actuator torque saturation protection.
[0044] In the present application, after detecting the actuator torque saturation, the process of attenuating the pilot control command is as follows: if the actuator torque is saturated for the first time, the control command is attenuated by a certain percentage, so as to reduce the aircraft motion response and reduce the demand for rudder surface augmentation; if the actuator torque is still saturated thereafter, the control command is attenuated by a certain percentage on the basis of the first attenuation; and so on, until the actuator torque saturation no longer occurs. Exemplarily, the attenuation percentage can be set to 30% or 50% and the like.
[0045] In the present application, after identifying the actuator torque saturation, the integrator is disconnected by setting the integrator on-off switch to the "off" state, so as to avoid integrator saturation; when detecting that the actuator torque is not saturated, the integrator is connected by setting the integrator on-off switch to the "on" state, so as to restore the normal control architecture.
[0046] On this basis, the present application further provides an actuator torque saturation protection control device for a statically unstable aircraft, which includes a forward control channel, a proportional control channel, an integral control channel, and an actuator torque saturation identification module, an integrator on-off switch module and a command attenuation module. The actuator torque saturation identification module is used to identify the actuator torque saturation and determine whether the actuator torque saturation occurs. The command attenuation module is used to attenuate the pilot control command when the actuator torque saturation occurs. The integrator on-off switch module is used to disconnect when the actuator torque saturation occurs, so as to achieve the disconnection of the integrator control channel.
[0047] The actuator torque saturation protection control method and device for a statically unstable aircraft provided by the present application can avoid the divergence of the aircraft caused by the actuator torque saturation of the statically unstable aircraft. The method does not need to modify any hardware part of the flight control system. The actuator torque saturation protection control law can be designed according to the method and added to the flight control law of the flight control system, so as to save the modification cost and achieve the protection of the actuator torque saturation of the statically unstable aircraft.
[0048] The above merely provides the specific implementation of the present application, but the protection scope of the present application is not limited to this. Any person skilled in the art can easily think of the changes or replacements within the technical range disclosed by the present application, which should be covered in the protection scope of the present application. Therefore, the protection scope of the present application should be subject to the protection scope of the claims.
Claims
1. A method for torque saturation protection control of statically unstable aircraft actuators, characterized in that, include: Identify actuator torque saturation and determine whether actuator torque saturation has occurred; When the actuator torque is saturated, an actuator torque saturation protection operation is executed. The actuator torque saturation protection operation includes attenuating the pilot's control commands and disconnecting the integrator control channel, thereby realizing the protection setting of the control law.
2. The torque saturation protection control method for statically unstable aircraft actuators as described in claim 1, characterized in that, The method for determining whether actuator torque saturation has occurred is as follows: a) The rudder surface position and the rudder deflection command have a predetermined deviation M for a continuous time period T; b) The speed at which the control surface moves in the direction of the rudder deflection command is lower than the predetermined speed threshold K, or the direction of the control surface moves in the opposite direction to the rudder deflection command; c) The dynamic pressure signal is greater than the dynamic pressure threshold S; If all three conditions are met, the actuator is judged to have torque saturation.
3. The torque saturation protection control method for statically unstable aircraft actuators as described in claim 2, characterized in that, The time period T is greater than the maximum response time of the actuator, which is the time it takes for the actuator to reach its maximum displacement from startup when the maximum displacement command is applied to the actuator.
4. The torque saturation protection control method for statically unstable aircraft actuators as described in claim 2, characterized in that, The deviation M is 5% to 10% of the rudder deflection command.
5. The torque saturation protection control method for statically unstable aircraft actuators as described in claim 2, characterized in that, The speed threshold K is the rudder deflection speed under the maximum hinge torque, which is determined based on the actuator loading test.
6. The torque saturation protection control method for statically unstable aircraft actuators as described in claim 2, characterized in that, The dynamic pressure threshold S is determined by using hinge torque data obtained from calculations and wind tunnel tests, as well as the output force performance of the actuator.
7. The torque saturation protection control method for statically unstable aircraft actuators as described in claim 1, characterized in that, When the actuator torque saturates, the process of attenuating the pilot's control commands is as follows: If the actuator experiences torque saturation for the first time, the pilot's control commands will be attenuated by a predetermined proportion, thereby reducing the aircraft's motion response and lowering the need for control surface stabilization. If torque saturation still occurs in the actuator, the pilot's control commands will be attenuated by a predetermined proportion on top of the first attenuation. This process continues until actuator torque saturation no longer occurs.
8. The torque saturation protection control method for statically unstable aircraft actuators as described in claim 7, characterized in that, The predetermined ratio is 30% to 50%.
9. The torque saturation protection control method for statically unstable aircraft actuators as described in claim 7, characterized in that, When the actuator torque is saturated, the method to disconnect the integrator control channel is as follows: By setting the integrator on / off switch to the "off" state, the integrator control channel is disconnected, thus preventing integrator saturation. When the actuator torque is not detected to be saturated, the integrator control channel is reconnected by setting the integrator on / off switch to the "on" state, restoring the normal control architecture.
10. An apparatus for implementing the torque saturation protection control method for statically unstable aircraft actuators as described in any one of claims 1-9, characterized in that, It includes a forward control channel, a proportional control channel, an integral control channel, an actuator torque saturation identification module, an integrator on / off switch module, and a command attenuation module. The actuator torque saturation identification module is used to identify actuator torque saturation and determine whether actuator torque saturation has occurred. The command attenuation module is used to attenuate the pilot's control commands when the actuator torque is saturated. The integrator on / off switch module is used to disconnect when the actuator torque is saturated, thereby disconnecting the integrator control channel.