Large aircraft configuration change compensation control method

Abstract

The invention provides a large aircraft configuration change compensation control method, and belongs to the technical field of flight control, and the method comprises the steps: S1, analyzing a flap, a slat and an air brake which can affect the aircraft configuration, and determining the aerodynamic characteristic change rule of the large aircraft when the flap, the slat and the air brake are retracted and released; s2, a flap configuration change compensation module and an air brake configuration change compensation module are constructed according to the aerodynamic characteristic change rule of the large aircraft when the flap, the slat and the air brake are retracted and released, and the flap configuration change compensation module and the air brake configuration change compensation module are arranged behind an integrator of a control law; therefore, the elevator can be adopted to balance pitching moment change at the beginning of aircraft configuration change, and then a configuration compensation control instruction of the elevator is replaced by the stabilizer.

Description

Technical Field

[0001] This application belongs to the field of flight control technology, and specifically relates to a method for compensating for changes in the configuration of large aircraft. Background Technology

[0002] To achieve optimal lift and drag, large aircraft typically feature flaps and speed brakes. The deployment and retraction of these flaps and speed brakes alter the aircraft's aerodynamic shape, also known as a flight configuration change, thus resulting in changes in aerodynamic characteristics. Deploying the flaps increases lift and generates a significant pitching moment, reducing the aircraft's pitch angle; deploying the speed brakes decreases lift and generates a significant pitching moment, increasing the aircraft's pitch angle.

[0003] Therefore, in order to reduce the pilot's operational burden, it is necessary to design a configuration change compensation control law or control method to suppress the pitch moment changes when the flaps and speed brakes are extended or retracted through the elevator and stabilizer, thereby reducing the pitch angle changes when the aircraft configuration changes. Summary of the Invention

[0004] The purpose of this application is to provide a large aircraft configuration change compensation control method to solve or mitigate at least one of the problems in the prior art.

[0005] The technical solution of this application is: a large aircraft configuration change compensation control method, including:

[0006] Step S1: Analyze the flaps, slats, and speed brakes that can affect the aircraft configuration, and determine the aerodynamic characteristics of the large aircraft when the flaps, slats, and speed brakes are extended or retracted.

[0007] Step S2: Based on the aerodynamic characteristic change law of the large aircraft when the flaps, slats and speed brakes are extended and retracted, construct the flap configuration change compensation module and the speed brake configuration change compensation module. The flap configuration change compensation module and the speed brake configuration change compensation module are set after the integrator of the control law, so that the elevator can be used to balance the pitch moment change at the beginning of the aircraft configuration change, and then the configuration compensation control command of the stabilizer is replaced by the elevator.

[0008] Preferably, the flap configuration change compensation module obtains the flap configuration change elevator compensation command by dividing the increase in pitch moment coefficient caused by the flap configuration change by the slope of the elevator pitch moment coefficient and passing it through the flap inertia link.

[0009] Preferably, the slope of the elevator pitch moment coefficient is determined based on aircraft wind tunnel test data.

[0010] Preferably, the time constant T of the inertial element is 1 / 3 of the movement time of each flap position.

[0011] Preferably, the speed brake configuration change compensation module divides the increase in pitch moment coefficient caused by the speed brake configuration change by the slope of the elevator pitch moment coefficient, and then obtains the speed brake configuration change elevator compensation command through the speed brake inertia link.

[0012] Preferably, the slope of the pitch moment coefficient of the speed brake is determined based on aircraft wind tunnel test data.

[0013] Preferably, the inertial element of the speed reducer is set to 1 / (0.5×s+1).

[0014] Preferably, the stabilizer motion command is generated by the stabilizer control logic module. The stabilizer is in discrete command control mode. When the command is 1, it means that the stabilizer controls the aircraft to tilt down and turn. When the command is -1, it means that the stabilizer controls the aircraft to tilt up and turn. When the command is 0, it means that the stabilizer remains in the current position.

[0015] Preferably, the process of generating the stabilized surface motion command is as follows:

[0016] 1) When the absolute value of the overload difference DNZ is less than 0.1:

[0017] a) When the absolute value of the integration command channel Int_cmd is greater than 2°, the motion condition of the stabilized surface is satisfied;

[0018] If the integration command channel Int_cmd≥0, the stabilizer motion command is 1; if the integration command channel Int_cmd<0, the stabilizer motion command is -1.

[0019] b) When the absolute value of the integration command channel Int_cmd is less than 0.5°, the stabilization motion condition is not met, and the stabilization motion command is 0.

[0020] c) When the absolute value of the integration command channel Int_cmd is greater than or equal to 0.5° and less than or equal to 2°, the stabilizer motion command retains the calculated value of the sampling step size from the previous time step.

[0021] 2) When the absolute value of the overload difference DNZ is greater than or equal to 0.1, the stabilizer motion command is 0.

[0022] The large aircraft configuration change compensation control method provided in this application suppresses changes in aircraft attitude and reduces the pilot's control burden by compensating for configuration changes when the flaps and speed brakes of large aircraft are extended and retracted. It does not require any modification to the hardware of the flight control system. The configuration change compensation 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 compensation control for configuration changes of the flaps and speed brakes of large aircraft. Attached Figure Description

[0023] 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.

[0024] Figure 1 This is a schematic diagram of a typical longitudinal control law architecture for a large aircraft.

[0025] Figure 2 This is a schematic diagram of the longitudinal control law architecture for a large aircraft according to this application.

[0026] Figure 3 This is a schematic diagram of the flap configuration change compensation module of this application.

[0027] Figure 4 This is a schematic diagram of the speed reducer configuration change compensation module of 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 addresses the problem of large pitch angle changes caused by the retraction and extension of flaps and speed brakes in large aircraft by proposing a configuration change compensation control method. This method compensates for pitch moment changes during the retraction and extension of flaps and speed brakes, thereby suppressing changes in aircraft attitude and reducing the pilot's control workload.

[0030] This application provides a method for compensating for changes in the configuration of large aircraft, including the following steps:

[0031] Step S1: Analyze the flaps, slats, and speed brakes that can affect the aircraft configuration to determine the aerodynamic characteristics of large aircraft when the flaps, slats, and speed brakes are extended or retracted.

[0032] To achieve optimal lift and drag, large aircraft typically incorporate flaps, slats, and speed brakes. The deployment and retraction of these components alter the aircraft's aerodynamic shape (also known as flight configuration changes), thus affecting its aerodynamic characteristics. Deploying flaps and slats increases lift and generates a significant pitching moment, reducing the aircraft's pitch angle; deploying speed brakes decreases lift and generates a significant pitching moment, increasing the aircraft's pitch angle.

[0033] To reduce the pilot's workload, it is necessary to use elevators and stabilizers for compensatory control to suppress pitch moment changes when flaps, slats, and speed brakes are extended or retracted, that is, to reduce pitch angle changes when the aircraft configuration changes.

[0034] The slats mainly change the inflection point of lift and pitch moment characteristics. When the slats are extended or retracted, they do not generate an increase in pitch moment. Therefore, when the flaps and slats are extended or retracted, only configuration change compensation control is needed for the flaps.

[0035] Step S2: Based on the aerodynamic characteristic change law of the large aircraft when the flaps, slats and speed brakes are extended and retracted, construct the flap configuration change compensation module and the speed brake configuration change compensation module. The flap configuration change compensation module and the speed brake configuration change compensation module are set after the integrator of the control law, so that the elevator can be used to balance the pitch moment change at the beginning of the aircraft configuration change, and then the configuration compensation control command of the stabilizer is replaced by the elevator.

[0036] like Figure 1 The diagram shows a typical longitudinal control law architecture for a large aircraft, consisting of a forward channel, a proportional channel, and an integral channel. The forward channel converts the control stick displacement into an overload command via a stick displacement-overload gradient module, then multiplies it by a gain k3 (the rudder deflection required per unit overload) to convert it into a forward rudder deflection command, enabling rapid aircraft response. The proportional channel introduces pitch rate and normal overload feedback to improve the dynamic response to normal overload, multiplying it by gains k1 and k2 respectively to convert it into a proportional rudder deflection command. The integral channel achieves precise control of the normal overload command through an integrator. It multiplies the overload difference (DMZ) between the actual normal overload and the normal overload command by a gain k4, and converts it into an integral rudder deflection command after passing through an integrator. The forward rudder deflection command, the proportional rudder deflection command, and the integral rudder deflection command together generate the elevator deflection command. Furthermore, the integral channel command Int_cmd generates a stabilizer command through the stabilizer control logic module, and after feedback by a gain k5, it cancels the value accumulated by the integrator, thus realizing the stabilizer replacing the elevator integral channel command.

[0037] The principle of flap configuration change compensation control and speed brake change compensation control is to initially use the elevator to balance the pitch moment change, and then use the stabilizer to replace the elevator configuration compensation control command, thereby avoiding occupying the elevator control authority. Based on this idea, in this application... Figure 1 Based on the control law architecture, a flap configuration change compensation module and a speed brake configuration change compensation module are added. These two modules are then integrated into the integrator, such as... Figure 2 As shown.

[0038] like Figure 3 As shown, the working principle of the flap configuration change compensation module of this application is as follows: the pitch moment coefficient increment generated by the flap configuration change is obtained according to the wind tunnel test data. After dividing it by the elevator efficiency (i.e. the slope of the pitch moment coefficient generated by the elevator), the angle at which the elevator balances the pitch moment can be obtained. After passing through the flap inertia link, the elevator compensation command for the flap configuration change can be obtained.

[0039] The pitch moment coefficient increment in this application is determined based on aircraft wind tunnel test data and is typically related to the flap handle position and angle of attack. Aircraft flaps have multiple positions, and the pilot controls the flaps to reach the target position by manipulating the flap handle. During aircraft wind tunnel testing, tests are conducted according to different flap configurations at each position, without testing every real-time position during flap movement. Therefore, the final wind tunnel test data obtained are data for each flap handle position, and an interpolation table of pitch moment coefficient increments (DM_FLAP) can be established based on the aerodynamic data, using flap handle position and angle of attack as variables.

[0040] In some embodiments of this application, the elevator pitch moment coefficient slope is determined based on aircraft wind tunnel test data and is typically related to flap position, slat position, Mach number, and angle of attack. An elevator pitch moment coefficient slope interpolation table (K_Mde) can be established based on aerodynamic data, using flap position, slat position, Mach number, and angle of attack as variables. Dividing the pitch moment coefficient increment interpolation table (DM_FLAP) by the elevator pitch moment coefficient slope interpolation table (K_Mde) yields the elevator compensation command for flap configuration changes.

[0041] However, since the flaps only move at a certain deflection speed after the pilot moves the flap handle to the desired position, and the flap movement stops after reaching the flap position command corresponding to the flap handle. When the flap handle changes, the pitch moment coefficient increment interpolation table DM_FLAP, established based on aerodynamic data, will abruptly change, directly generating the pitch moment coefficient increment for the target flap position. Therefore, considering that the pitch moment generated during the flap configuration change is gradually changing, this application sets a flap inertial element module after dividing the pitch moment coefficient increment by the elevator pitch moment coefficient slope, and its time constant T is determined according to the flap movement time. According to the fact that the time required for the flap inertial element to reach a steady state value is 3 times T (the 3 times T principle), the time constant T is the movement time of each flap position divided by 3. If the time for each position is different, the time constant T changes according to the flap handle position.

[0042] like Figure 4 As shown, the working principle of the speed brake configuration change compensation module in this application is as follows: the pitch moment coefficient increment caused by the speed brake configuration change can be obtained based on wind tunnel test data and the real-time angle of the speed brake. After dividing by the elevator efficiency, the angle at which the elevator balances the pitch moment can be obtained. After passing through the speed brake inertia link, the elevator compensation command for the speed brake configuration change can be obtained.

[0043] In some embodiments of this application, the slope of the speed brake pitch moment coefficient is determined based on aircraft wind tunnel test data, and is typically related to the flap position, slat position, Mach number, and angle of attack. An interpolation table (K_Msp) of the speed brake pitch moment coefficient slope can be established based on aerodynamic data, with flap position, slat position, Mach number, and angle of attack as variables. Multiplying the speed brake pitch moment coefficient slope interpolation table (K_Msp) by the speed brake angle and then dividing by the elevator pitch moment coefficient slope interpolation table (K_Mde) yields the elevator compensation command for speed brake configuration changes.

[0044] In this application, in order to make the compensation command for the configuration change caused by the opening of the speed brake smoother, a speed brake inertial element is set after dividing by the elevator efficiency. The speed brake inertial element can be set to 1 / (0.5×s+1).

[0045] The control process of the stabilizer control logic module in the control law of this application is as follows: when the integral channel command Int_cmd is greater than a certain threshold and the aircraft is in steady-state flight, a stabilizer motion command is generated according to the polarity of the integral command channel Int_cmd. The aircraft is in steady-state flight according to... Figure 2 The overload difference DNZ is used for judgment. The stabilizer is in discrete command control mode, that is, when the command is 1, it means that the stabilizer controls the aircraft to pitch down and yaw, when the command is -1, it means that the stabilizer controls the aircraft to pitch up and yaw, and when the command is 0, it means that the stabilizer remains in the current position.

[0046] The specific logic of the stabilizer control command is as follows:

[0047] 1) When the absolute value of the overload difference DNZ is less than 0.1:

[0048] a) When the absolute value of the integral instruction channel Int_cmd is greater than 2°, the motion condition of the stabilizer is satisfied.

[0049] If the integration command channel Int_cmd≥0, the stabilizer motion command is 1; if the integration command channel Int_cmd<0, the stabilizer motion command is -1.

[0050] b) When the absolute value of the integration command channel Int_cmd is less than 0.5°, the stabilization motion condition is not met, and the stabilization motion command is 0.

[0051] c) When the absolute value of the integration command channel Int_cmd is greater than or equal to 0.5° and less than or equal to 2°, the stabilizer motion command retains the calculated value of the sampling step size from the previous step.

[0052] 2) When the absolute value of the overload difference DNZ is greater than or equal to 0.1, the stabilizer motion command is 0.

[0053] The stick displacement-overload gradient module in the control law of this application is designed based on the aircraft's normal overload limit and the stick force overload gradient requirements in flight quality requirements. The gains k1, k2, k3, and k4 are determined based on the normal overload response to ensure fast initial response, good dynamic response, and high control accuracy.

[0054] In this application, the gain k5 in the control law is used to cancel the elevator integral channel command when the stabilizer moves. In order to ensure a smooth handover of control between the stabilizer and the elevator, k5 is the ratio of the efficiency of the pitching moment generated by the elevator and the stabilizer, which can be calculated based on the aircraft wind tunnel data.

[0055] The large aircraft configuration change compensation control method provided in this application suppresses changes in aircraft attitude and reduces the pilot's control burden by compensating for configuration changes when the flaps and speed brakes of large aircraft are extended and retracted. It does not require any modification to the hardware of the flight control system. The configuration change compensation 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 compensation control for configuration changes of the flaps and speed brakes of large aircraft.

[0056] 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 compensating for changes in the configuration of a large aircraft, characterized in that, include: Step S1: Analyze the flaps, slats, and speed brakes that can affect the aircraft configuration, and determine the aerodynamic characteristics of the large aircraft when the flaps, slats, and speed brakes are extended or retracted. Step S2: Based on the aerodynamic characteristic change law of the large aircraft when the flaps, slats and speed brakes are extended and retracted, construct the flap configuration change compensation module and the speed brake configuration change compensation module. The flap configuration change compensation module and the speed brake configuration change compensation module are set after the integrator of the control law, so that the elevator can be used to balance the pitch moment change at the beginning of the aircraft configuration change, and then the configuration compensation control command of the stabilizer is replaced by the elevator.

2. The large aircraft configuration change compensation control method as described in claim 1, characterized in that, The flap configuration change compensation module divides the increase in pitch moment coefficient caused by the flap configuration change by the slope of the elevator pitch moment coefficient, and then passes through the flap inertia link to obtain the flap configuration change elevator compensation command.

3. The large aircraft configuration change compensation control method as described in claim 2, characterized in that, The slope of the elevator pitch moment coefficient is determined based on aircraft wind tunnel test data.

4. The large aircraft configuration change compensation control method as described in claim 2, characterized in that, The time constant T of the inertial element is 1 / 3 of the movement time of each flap position.

5. The large aircraft configuration change compensation control method as described in any one of claims 2 to 4, characterized in that, The speed brake configuration change compensation module divides the increase in pitch moment coefficient caused by the speed brake configuration change by the slope of the elevator pitch moment coefficient, and then obtains the speed brake configuration change elevator compensation command through the speed brake inertia link.

6. The large aircraft configuration change compensation control method as described in claim 5, characterized in that, The slope of the pitch moment coefficient of the speed brake is determined based on aircraft wind tunnel test data.

7. The large aircraft configuration change compensation control method as described in claim 5, characterized in that, The inertial element of the speed reducer is set to 1 / (0.5×s+1).

8. The large aircraft configuration change compensation control method as described in claim 5, characterized in that, The stabilizer control logic module generates stabilizer motion commands. The stabilizer is in discrete command control mode. When the command is 1, it means that the stabilizer controls the aircraft to pitch down and yaw. When the command is -1, it means that the stabilizer controls the aircraft to pitch up and yaw. When the command is 0, it means that the stabilizer remains in its current position.

9. The large aircraft configuration change compensation control method as described in claim 8, characterized in that, The process of generating the stabilized surface motion command is as follows: 1) When the absolute value of the overload difference DNZ is less than 0.1: a) When the absolute value of the integration command channel Int_cmd is greater than 2°, the motion condition of the stabilized surface is satisfied; If the integration command channel Int_cmd≥0, the stabilizer motion command is 1; if the integration command channel Int_cmd<0, the stabilizer motion command is -1. b) When the absolute value of the integration command channel Int_cmd is less than 0.5°, the stabilization motion condition is not met, and the stabilization motion command is 0. c) When the absolute value of the integration command channel Int_cmd is greater than or equal to 0.5° and less than or equal to 2°, the stabilizer motion command retains the calculated value of the sampling step size from the previous time step. 2) When the absolute value of the overload difference DNZ is greater than or equal to 0.1, the stabilizer motion command is 0.

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