Flap configuration change compensation control method and system for large amphibious aircraft
By analyzing the changes in aerodynamic lift coefficient caused by flap angle changes, an angle of attack increment compensation value is generated and filtered to generate compensation commands when the flap angle changes. This solves the pilot's operational burden problem caused by flap changes in large amphibious aircraft, and improves flight safety and piloting experience.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- XIAN FLIGHT SELF CONTROL INST OF AVIC
- Filing Date
- 2024-02-23
- Publication Date
- 2026-06-09
Smart Images
Figure CN118062225B_ABST
Abstract
Description
Technical Field
[0001] This invention relates to, but is not limited to, the field of flight control system technology, and particularly to a method and system for compensating for flap configuration changes in a large amphibious aircraft. Background Technology
[0002] To reduce the water load, large amphibious aircraft typically use retractable flaps to slow down the aircraft's landing speed.
[0003] During flight, the retractable flaps significantly alter the aircraft's aerodynamic characteristics, causing noticeable changes in the forces and moments acting on the aircraft. This results in a more abrupt pitching up or down, increasing the pilot's workload and negatively impacting the flight experience. Summary of the Invention
[0004] The purpose of this invention is to solve the above-mentioned technical problems. This invention provides a method and system for compensating for flap configuration changes in large amphibious aircraft. This addresses the problems that existing large amphibious aircraft, due to the use of retractable flaps, experience significant changes in their aerodynamic characteristics during flight, leading to a sharp pitching or plunging tendency, increasing the pilot's workload, and negatively impacting the pilot's experience.
[0005] The technical solution of the present invention: The embodiments of the present invention provide a method for compensating for flap configuration changes in a large amphibious aircraft, comprising:
[0006] Step 1: Obtain the relationship between the aerodynamic lift coefficient and the angle of attack for the aircraft at different flap angles;
[0007] Step 2: Select the appropriate flap signal based on whether the flap is malfunctioning;
[0008] Step 3: Using the selected flap signal as the input variable, interpolation is performed based on the relationship between the aerodynamic lift coefficient and the angle of attack to obtain the angle of attack increment compensation value. By performing high-pass filtering and amplitude limiting on the angle of attack increment compensation value and multiplying it by different gains, the compensation commands for different longitudinal branches when the flap angle changes are obtained.
[0009] Step 4: Connect the compensation commands for different longitudinal branches when the flap angle changes to the corresponding longitudinal control law to achieve the compensation function when the flap angle changes.
[0010] Optionally, in the flap configuration change compensation control method for large amphibious aircraft as described above, step 1 includes:
[0011] Based on the aircraft configuration and lift coefficient curves under different flap angles, the relationship between the aerodynamic lift coefficient and the angle of attack under different flap angles is analyzed.
[0012] Optionally, in the flap configuration change compensation control method for large amphibious aircraft as described above, step 2 includes:
[0013] The value of the flap signal is determined based on whether the flap is malfunctioning;
[0014] When the flap signal is normal, select the current flap signal; when the flap signal is faulty, select the previous flap signal.
[0015] Optionally, in the flap configuration change compensation control method for large amphibious aircraft as described above, the compensation commands for different longitudinal branches when the flap angle changes in step 3 include: flap angle change pitch rate compensation command and flap angle change pitch rate increment command.
[0016] Optionally, in the flap configuration change compensation control method for large amphibious aircraft as described above, step 3 includes:
[0017] Step 31: Using the selected flap signal as the input variable, obtain the change in angle of attack during different flap angle changes, and form an interpolation table of constant lift coefficient angle of attack increment.
[0018] Step 32: Use the constant lift coefficient angle of attack increment interpolation table to obtain the angle of attack increment compensation value under different flap angles;
[0019] Step 33: Pass the angle of attack increment compensation value command through a high-pass filter to obtain the pitch rate compensation command required for the angle of attack increment when the flap angle changes during the flap retraction and extension process.
[0020] Step 34: After passing the pitch rate compensation command through the limiting circuit and multiplying it by the gain A, the pitch rate compensation command for flap angle change is obtained.
[0021] Step 35: After passing the pitch rate compensation command through the limiting stage and multiplying it by the gain B, the pitch rate increment command for flap angle change is obtained.
[0022] Optionally, in the flap configuration change compensation control method for large amphibious aircraft as described above, step 31, which involves forming an interpolation table of constant lift coefficient angle of attack increments, includes:
[0023] Based on the aircraft lift coefficient curve, within the linear region of the basic lift coefficient, a constant lift coefficient straight line is determined for different flap angles. Based on the constant lift coefficient straight line, the change in angle of attack during the change of different flap angles is confirmed, forming a constant lift coefficient angle of attack increment interpolation table.
[0024] Optionally, in the flap configuration change compensation control method for large amphibious aircraft described above, the specific method for forming the constant lift coefficient angle of attack increment interpolation table in step 31 is as follows:
[0025] Step 31a: For the lift coefficient curve, within the linear region of the basic lift coefficient, select the 0-degree flap lift coefficient curve corresponding to the specified angle of attack and draw a straight line with equal lift coefficient.
[0026] Step 31b: Obtain the angle of attack value for each flap angle under the given constant lift coefficient straight line;
[0027] Step 31c: Using the 0-degree flap as a reference, obtain the change in angle of attack corresponding to the transition from 0 degrees flap to each flap angle, which is the angle of attack increment compensation value under different flap angles.
[0028] The specified angle of attack is selected within the range of angles of attack during a normal flap retraction / extension scenario.
[0029] Optionally, in the flap configuration change compensation control method for large amphibious aircraft as described above, step 4 includes:
[0030] The flap angle change pitch rate increment command is superimposed with the aircraft pitch rate signal and then enters the integral control loop.
[0031] The flap configuration change pitch rate compensation command is superimposed on the aircraft pitch rate signal and then enters the damping control loop.
[0032] This invention also provides a flap configuration change compensation control system for a large amphibious aircraft. The flap angle change compensation control system is used to execute the flap configuration change compensation control method for a large amphibious aircraft as described in any of the above claims. The control includes: a lift coefficient relationship generation module, a flap signal selection module, a compensation command generation module, a compensation command output module, an integral control loop, and a damping control loop.
[0033] The lift coefficient relationship generation module is used to obtain the relationship between the aerodynamic lift coefficient and the angle of attack of an aircraft at different flap angles;
[0034] The flap signal selection module is used to select the appropriate flap signal based on whether the flap is malfunctioning.
[0035] The compensation command generation module takes the flap signal selected by the flap signal selection module as the input variable, and interpolates the aerodynamic lift coefficient with the angle of attack obtained by the lift coefficient relationship generation module to obtain the angle of attack increment compensation value. By performing high-pass filtering and amplitude limiting on the angle of attack increment compensation value and multiplying it by different gains, the compensation commands for different longitudinal branches when the flap angle changes are obtained. The compensation commands for different longitudinal branches when the flap angle changes include: the flap angle change pitch rate compensation command and the flap angle change pitch rate increment command.
[0036] The compensation command output module is used to input the compensation commands of different longitudinal branches when the flap angle changes into the corresponding longitudinal control law to realize the compensation function when the flap angle changes. Specifically, the flap angle change pitch rate increment command is superimposed with the aircraft pitch rate signal and then enters the integral control loop; the flap configuration change pitch rate compensation command is superimposed with the aircraft pitch rate signal and then enters the damping control loop.
[0037] Optionally, in the flap configuration change compensation control system of a large amphibious aircraft as described above, the compensation command generation module specifically includes:
[0038] The interpolation table unit is used to obtain the change in angle of attack during different flap angle changes by using the selected flap signal as the input variable, and to form a constant lift coefficient angle of attack increment interpolation table; it is also used to obtain the angle of attack increment compensation value under different flap angles using the constant lift coefficient angle of attack increment interpolation table.
[0039] The filtering unit is used to pass the angle of attack increment compensation value command through a high-pass filter to obtain the pitch rate compensation command required for the angle of attack increment when the flap angle changes during the flap retraction and extension process.
[0040] The limiting unit is used to pass the pitch rate compensation command through the limiting circuit and multiply it by gain A to obtain the flap angle change pitch rate compensation command; it is also used to pass the pitch rate compensation command through the limiting circuit and multiply it by gain B to obtain the flap angle change pitch rate increment command.
[0041] The beneficial effects of this invention: This invention provides a method and system for compensating for flap configuration changes in large amphibious aircraft. It obtains the relationship between the aerodynamic lift coefficient and the angle of attack at different flap angles; selects the appropriate flap signal based on whether the flap is faulty; uses the selected flap signal as an input variable, interpolates the relationship between the aerodynamic lift coefficient and the angle of attack to obtain an angle-of-attack increment compensation value; performs high-pass filtering and amplitude limiting on the angle-of-attack increment compensation value, and multiplies it by different gains to obtain compensation commands for different longitudinal branches when the flap angle changes; and integrates the compensation commands for different longitudinal branches when the flap angle changes into the corresponding longitudinal control law to achieve the compensation function when the flap angle changes. By adopting the technical solution of this invention, the adverse effects generated during flap deployment and retraction are automatically compensated, reducing the pilot's operational burden and improving flight safety to a certain extent. The beneficial effects of this invention are explained below:
[0042] First, a flap configuration compensation control strategy based on aerodynamic lift coefficient is adopted. The technical solution of this invention analyzes the change in aerodynamic lift coefficient caused by the change in flap angle. In the linear region of the basic quantity of aerodynamic lift coefficient, the increment of angle of attack under different flaps is identified through the constant lift coefficient line, thereby forming an interpolation table. After high-pass filtering and amplitude limiting, and multiplied by the relevant gain, it is superimposed on the damping branch and integration branch of the longitudinal control structure to compensate the pitch angular velocity signal, realize compensation control when the flap angle changes, and reduce the pilot's operational burden when the flap angle changes.
[0043] Second, a dual-loop compensation strategy is adopted, consisting of an integral control loop and a damping control loop for precise angular velocity control. The dual-loop compensation technology generates different compensation commands at different flap angles and superimposes these commands onto the integral control loop and damping control loop of the main flight controller, respectively, to achieve compensation control when the flaps change. This reduces the pilot's workload to some extent and improves the flight experience.
[0044] Third, a fault-tolerant protection strategy is adopted by using flap configuration to compensate for the failure.
[0045] On the one hand, it realizes flap signal fault protection: when the flap signal fails, the previous flap signal is taken as the flap input to ensure that no additional damage is caused to the aircraft; when the flap signal is normal, the current flap angle is selected.
[0046] On the other hand, it achieves amplitude limiting safety protection: through the amplitude limiting link, it ensures that when there is a constant input, the compensation control method has a small impact on the aircraft operation, and does not affect the normal take-off and landing of the aircraft.
[0047] Furthermore, the technical solution provided in this embodiment of the invention has been verified in a simulator. Based on the actual response results of the aircraft, the flap angle change compensation control method provided in this embodiment of the invention has a significant effect on improving the piloting experience of large amphibious aircraft and reducing the pilot's operational burden. Attached Figure Description
[0048] The accompanying drawings are provided to further understand the technical solutions of the present invention and constitute a part of the specification. They are used together with the embodiments of this application to explain the technical solutions of the present invention and do not constitute a limitation on the technical solutions of the present invention.
[0049] Figure 1 A flowchart of a flap configuration change compensation control method for a large amphibious aircraft provided in an embodiment of the present invention;
[0050] Figure 2 This is a schematic diagram of the control principle of a flap configuration change compensation control method for a large amphibious aircraft provided in an embodiment of the present invention;
[0051] Figure 3 This is a schematic diagram illustrating the confirmation of the angle of attack increment interpolation table in the flap angle change compensation control method provided in this embodiment of the invention;
[0052] Figure 4 This is a schematic diagram of the simulation results of flap angle change compensation control in Embodiment 1 of the present invention;
[0053] Figure 5 This is a schematic diagram of the simulation results of flap angle change compensation control in Embodiment 2 of the present invention. Detailed Implementation
[0054] To make the objectives, technical solutions, and advantages of the present invention clearer, the embodiments of the present invention will be described in detail below with reference to the accompanying drawings. It should be noted that, unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.
[0055] As explained in the background section, existing large amphibious aircraft, due to the use of retractable flaps, experience significant changes in their aerodynamic characteristics during flight. This results in a more abrupt pitching up or down tendency, increasing the pilot's workload and negatively impacting the piloting experience.
[0056] To address the aforementioned problems, this invention provides a flap configuration change compensation control method and system for large amphibious aircraft. Based on the angle-of-attack increment of the constant lift coefficient at different flap angles, the flap retraction / extension speed is converted into a pitch rate to obtain the flap retraction / extension compensation pitch rate. This flap angle change compensation control method automatically controls the aircraft, compensating for changes in flap angle while maintaining the original flight altitude and ensuring stable flight. This effectively reduces the pilot's workload, improves the piloting experience, and guarantees flight safety.
[0057] The present invention provides the following specific embodiments, which can be combined with each other. For the same or similar concepts or processes, they may not be described again in some embodiments.
[0058] Figure 1 This is a flowchart illustrating a flap configuration change compensation control method for a large amphibious aircraft, provided by an embodiment of the present invention. Figure 2 This is a schematic diagram illustrating the control principle of a flap configuration change compensation control method for a large amphibious aircraft provided in an embodiment of the present invention.
[0059] The flap configuration change compensation control method for large amphibious aircraft provided in this embodiment of the invention may include the following steps:
[0060] Step 1: Obtain the relationship between the aerodynamic lift coefficient and the angle of attack for the aircraft at different flap angles;
[0061] In this step, the specific method for obtaining the relationship between the aerodynamic lift coefficient and the angle of attack is as follows: based on the aircraft configuration and the aircraft lift coefficient curve under different flap angles, the relationship between the aerodynamic lift coefficient and the angle of attack under different flap angles is analyzed.
[0062] Step 2: Select the appropriate flap signal based on whether the flap is malfunctioning;
[0063] In this step, the specific method for selecting the appropriate flap signal is as follows: when the flap signal is normal, the current flap signal is selected; when the flap signal malfunctions, the previous flap signal is selected, thereby ensuring that the flap signal does not change abruptly and ensuring flight safety.
[0064] Step 3: Using the selected flap signal as the input variable, interpolate to obtain the angle of attack increment compensation value. By performing high-pass filtering and amplitude limiting on the angle of attack increment compensation value and multiplying it by different gains, the compensation commands for different longitudinal branches when the flap angle changes are obtained.
[0065] Step 4: Connect the compensation commands for different longitudinal branches when the flap angle changes to the corresponding longitudinal control law to achieve the compensation function when the flap angle changes.
[0066] In this step, the compensation commands for different longitudinal branches when the flap angle changes include: the flap angle change pitch rate compensation command and the flap angle change pitch rate increment command.
[0067] The flap configuration change compensation control method for large amphibious aircraft provided in this embodiment of the invention can offset the adverse effects of configuration changes by compensating for pitch rate changes when flap angles change, thereby reducing the pilot's workload and improving safety to a certain extent.
[0068] In one implementation of this invention, step 3 may include:
[0069] Step 31: Using the selected flap signal as the input variable, obtain the change in angle of attack during different flap angle changes, and form an interpolation table of constant lift coefficient angle of attack increment.
[0070] The method for forming the constant lift coefficient angle of attack increment interpolation table in this step is as follows: Based on the aircraft lift coefficient curve (i.e., aerodynamic relationship data), within the linear region of the basic lift coefficient (meaning the lift coefficient only changes with the angle of attack), determine the constant lift coefficient straight line under different flap angles, and confirm the change in angle of attack during the change of different flap angles based on the constant lift coefficient straight line, thus forming the constant lift coefficient angle of attack increment interpolation table.
[0071] Step 32: Use the constant lift coefficient angle of attack increment interpolation table to obtain the angle of attack increment compensation value under different flap angles;
[0072] Step 33: Pass the angle of attack increment compensation value command through a high-pass filter to obtain the pitch rate compensation command required for the angle of attack increment when the flap angle changes during the flap retraction and extension process.
[0073] Step 34: After passing the pitch rate compensation command through the limiting circuit and multiplying it by the gain A, the pitch rate compensation command for flap angle change is obtained.
[0074] Step 35: After passing the pitch rate compensation command through the limiting stage and multiplying it by the gain B, the pitch rate increment command for flap angle change is obtained.
[0075] Furthermore, in this implementation method, the specific method for forming the constant lift coefficient angle of attack increment interpolation table in step 31 is as follows:
[0076] Step 31a: For the lift coefficient curve, within the linear region of the basic lift coefficient, select the 0-degree flap lift coefficient curve corresponding to the specified angle of attack and draw a straight line with equal lift coefficient.
[0077] Step 31b: Obtain the angle of attack value for each flap angle under the given constant lift coefficient straight line;
[0078] Step 31c: Using the 0-degree flap as a reference, obtain the change in angle of attack corresponding to the transition from 0 degrees flap to each flap angle, which is the angle of attack increment compensation value under different flap angles.
[0079] The specified angle of attack is selected within the range of angles of attack during normal flap retraction / extension.
[0080] In one implementation of this invention, step 4 described above may include:
[0081] On the one hand, the flap angle change pitch rate increment command is superimposed with the aircraft pitch rate signal and then enters the integral control loop;
[0082] On the other hand, the flap configuration change pitch rate compensation command is superimposed with the aircraft pitch rate signal and then enters the damping control loop.
[0083] In this implementation, the pitch rate is automatically compensated in the integral control loop and the damping control loop, which effectively suppresses the adverse response when the flap angle changes, and realizes the control function of flap configuration change compensation.
[0084] Based on the flap configuration change compensation control method for large amphibious aircraft provided in the embodiments of the present invention, the embodiments of the present invention also provide a flap configuration change compensation control system for large amphibious aircraft. The flap angle change compensation control system is used to execute the flap configuration change compensation control method for large amphibious aircraft in any of the above embodiments. The control system includes: a lift coefficient relationship generation module, a flap signal selection module, a compensation command generation module, a compensation command output module, as well as an integral control loop and a damping control loop.
[0085] See Figure 1 The control principle shown in this embodiment of the invention is that the lift coefficient relationship generation module is used to obtain the relationship between the aerodynamic lift coefficient and the angle of attack of the aircraft at different flap angles.
[0086] The flap signal selection module in this embodiment of the invention is used to select the corresponding flap signal according to whether the flap is malfunctioning.
[0087] The compensation command generation module in this embodiment of the invention is used to take the flap signal selected by the flap signal selection module as the input variable, and interpolate the aerodynamic lift coefficient with the angle of attack obtained by the lift coefficient relationship generation module to obtain the angle of attack increment compensation value. By performing high-pass filtering and amplitude limiting on the angle of attack increment compensation value, and multiplying it by different gains, the compensation commands for different longitudinal branches when the flap angle changes are obtained. The compensation commands for different longitudinal branches when the flap angle changes include: the flap angle change pitch rate compensation command and the flap angle change pitch rate increment command.
[0088] The compensation command output module in this embodiment of the invention is used to connect the compensation commands of different longitudinal branches when the flap angle changes to the corresponding longitudinal control law, so as to realize the compensation function when the flap angle changes; specifically, the flap angle change pitch rate increment command is superimposed with the aircraft pitch rate signal and then enters the integral control loop; the flap configuration change pitch rate compensation command is superimposed with the aircraft pitch rate signal and then enters the damping control loop.
[0089] In this embodiment of the invention, the compensation instruction generation module may specifically include:
[0090] The interpolation table unit is used to obtain the change in angle of attack during different flap angle changes by using the selected flap signal as the input variable, and to form a constant lift coefficient angle of attack increment interpolation table; it is also used to obtain the angle of attack increment compensation value under different flap angles using the constant lift coefficient angle of attack increment interpolation table.
[0091] In the specific implementation process, the method by which the interpolation table unit forms the angle of attack increment interpolation table of constant lift coefficient has been described in detail in the above embodiments, so it will not be repeated here.
[0092] The filtering unit is used to pass the angle of attack increment compensation value command through a high-pass filter to obtain the pitch rate compensation command required for the angle of attack increment when the flap angle changes during the flap retraction and extension process.
[0093] The limiting unit is used to pass the pitch rate compensation command through the limiting circuit and multiply it by gain A to obtain the flap angle change pitch rate compensation command; it is also used to pass the pitch rate compensation command through the limiting circuit and multiply it by gain B to obtain the flap angle change pitch rate increment command.
[0094] This invention provides a method and system for compensating for flap configuration changes in a large amphibious aircraft. The method involves acquiring the relationship between the aerodynamic lift coefficient and the angle of attack at different flap angles; selecting a corresponding flap signal based on whether the flap is malfunctioning; using the selected flap signal as an input variable, interpolating the relationship between the aerodynamic lift coefficient and the angle of attack to obtain an angle-of-attack increment compensation value; high-pass filtering and limiting the angle-of-attack increment compensation value, and multiplying it by different gains to obtain compensation commands for different longitudinal branches when the flap angle changes; and then integrating these compensation commands into the corresponding longitudinal control laws to achieve the compensation function when the flap angle changes. This invention's technical solution automatically compensates for adverse effects during flap deployment and retraction, reducing the pilot's workload and improving flight safety to some extent. The beneficial effects of this invention are described below:
[0095] First, a flap configuration compensation control strategy based on aerodynamic lift coefficient is adopted. The technical solution of this invention analyzes the change in aerodynamic lift coefficient caused by the change in flap angle. In the linear region of the basic quantity of aerodynamic lift coefficient, the increment of angle of attack under different flaps is identified through the constant lift coefficient line, thereby forming an interpolation table. After high-pass filtering and amplitude limiting, and multiplied by the relevant gain, it is superimposed on the damping branch and integration branch of the longitudinal control structure to compensate the pitch angular velocity signal, realize compensation control when the flap angle changes, and reduce the pilot's operational burden when the flap angle changes.
[0096] Second, a dual-loop compensation strategy is adopted, consisting of an integral control loop and a damping control loop for precise angular velocity control. The dual-loop compensation technology generates different compensation commands at different flap angles and superimposes these commands onto the integral control loop and damping control loop of the main flight controller, respectively, to achieve compensation control when the flaps change. This reduces the pilot's workload to some extent and improves the flight experience.
[0097] Third, a fault-tolerant protection strategy is adopted by using flap configuration to compensate for the failure.
[0098] On the one hand, it realizes flap signal fault protection: when the flap signal fails, the previous flap signal is taken as the flap input to ensure that no additional damage is caused to the aircraft; when the flap signal is normal, the current flap angle is selected.
[0099] On the other hand, it achieves amplitude limiting safety protection: through the amplitude limiting link, it ensures that when there is a constant input, the compensation control method has a small impact on the aircraft operation, and does not affect the normal take-off and landing of the aircraft.
[0100] Furthermore, the technical solution provided in this embodiment of the invention has been verified in a simulator. Based on the actual response results of the aircraft, the flap angle change compensation control method provided in this embodiment of the invention has a significant effect on improving the piloting experience of large amphibious aircraft and reducing the pilot's operational burden.
[0101] The following specific embodiments illustrate the implementation of the flap configuration change compensation control method and system for large amphibious aircraft provided by the present invention.
[0102] See Figure 1 The schematic diagram shown illustrates the principle of the flap angle change compensation control method. The flap angle change compensation control method provided in this embodiment of the invention includes the following steps:
[0103] Step 1: Aerodynamic Data Analysis
[0104] 1. Based on the aircraft configuration and lift coefficient curves under different flap angles, the relationship between the aerodynamic lift coefficient and the angle of attack under different flap angles is analyzed.
[0105] Step 2: Flap Signal Selection
[0106] 1. Determine the flap signal value based on whether the flap is malfunctioning. When the flap signal is normal, select the current flap signal. When the flap signal is malfunctioning, select the previous flap signal to ensure that the flap signal does not change abruptly and to ensure flight safety.
[0107] Step 3: Generate compensation instructions
[0108] In this step, based on the flap signal, the angle of attack increment compensation value under different flap angles is obtained by interpolating the equal lift coefficient angle of attack increment at different flap angles.
[0109] In this step, the angle of attack increment interpolation table is first obtained. Figure 3 This is a schematic diagram illustrating the confirmation of the angle-of-attack increment interpolation table in the flap angle change compensation control method provided in this embodiment of the invention. The angle-of-attack increment interpolation table is confirmed according to the following steps:
[0110] 1) According to Figure 3 As shown, by analyzing the deployment and retraction scenarios at different flap angles, the linear region of the basic lift coefficient was selected, and the straight line of constant lift coefficient was confirmed.
[0111] 2) According to Figure 3 The curve drawn in the figure shows that when the angle of attack is 6°, a straight line with the same lift coefficient is drawn based on the lift coefficient of the flap at 0° at this moment. Then, the angle of attack with the same lift coefficient of the flap at 17° is 1°, the angle of attack with the flap at 25° is generally -2°, and the angle of attack with the flap at 40° is generally -5°.
[0112] 3) Using 0 degrees of flap as a baseline, the angle of attack change Δ1 is -5° when transitioning from 0 degrees to 17 degrees, Δ2 is -8° when transitioning from 0 degrees to 25 degrees, and Δ3 is -11° when transitioning from 0 degrees to 40 degrees. This forms a constant lift coefficient angle of attack increment interpolation table. Using this table, the angle of attack increment compensation value for different flap angles can be obtained.
[0113] Secondly, the angle of attack increment compensation value command obtained by interpolation at different flap angles is passed through a high-pass filter to obtain the pitch rate compensation command required for the angle of attack increment of the aircraft angle of attack to the constant lift coefficient line at other flap angles during flap retraction and extension. This signal is matched with the flap retraction and extension rate and can reflect the flap change rate in real time, reducing the response when the flap configuration changes.
[0114] Furthermore, on the one hand, the pitch rate compensation command is passed through a limiting stage and multiplied by gain A to obtain the flap configuration change pitch rate compensation command; on the other hand, the pitch rate compensation command is passed through a limiting stage and multiplied by gain B to obtain the flap configuration change pitch rate increment command.
[0115] Step 4: Instruction Synthesis
[0116] 1. Maintain the original control architecture, add the flap configuration change pitch rate compensation command to the pitch rate signal to obtain the integrated damping branch command, and superimpose it into the damping control loop in the main flight control.
[0117] 2. Maintain the original control architecture, add the pitch rate increment command for flap configuration change to the pitch rate signal to obtain the integrated branch command, and superimpose it on the integral control loop in the main flight controller.
[0118] The flap angle change compensation control method provided in this embodiment of the invention can automatically compensate for pitch rate, effectively suppress adverse responses when flap angle changes, and realize the control function of flap configuration change compensation, thereby realizing the flap configuration change compensation control method for large amphibious aircraft mentioned in this invention.
[0119] Example 1
[0120] Step 1: Aerodynamic Data Analysis
[0121] In practice, based on the aircraft configuration and lift coefficient curves under different flap angles, the relationship between the aerodynamic lift coefficient and the angle of attack under different flap angles is analyzed.
[0122] Step 2: Flap Signal Selection
[0123] In practice, the initial flap angle is set to 0°, the flap fault signal is set to 0 (1 for fault, 0 for normal), the total simulation time is 15s, and the flap angle changes from 0° to 17° in 5s.
[0124] Step 3: Generate compensation instructions
[0125] 301. In specific implementation, based on aerodynamic coefficient analysis, an interpolation table of angle of attack increments under different flaps is obtained, from... Figure 3 The interpolation table is as follows;
[0126] Flaps (°) Angle of attack increment (°) 0 0 17 -5 25 -8 40 -11
[0127] 302. In specific implementation, the interpolated angle of attack increment signal is passed through a high-pass filter to obtain the real-time pitch rate compensation command;
[0128] 303. In specific implementation, the pitch rate compensation command is limited and then multiplied by gain A to obtain the flap configuration change pitch rate compensation command;
[0129] 304. In specific implementation, the pitch rate compensation command is limited and then multiplied by the gain B to obtain the pitch rate increment command for flap configuration change;
[0130] Step 4: Instruction Synthesis
[0131] The flap configuration change pitch rate compensation command and the flap configuration change pitch rate increment command are respectively integrated into the longitudinal damping control loop and integral control loop of the main flight control system, thereby realizing compensation control during flap configuration changes. Figure 4 The figure shown is a schematic diagram of the simulation results of flap angle change compensation control in Embodiment 1 of the present invention.
[0132] Example 2
[0133] Step 1: Aerodynamic Data Analysis
[0134] In practice, based on the aircraft configuration and lift coefficient curves under different flap angles, the relationship between the aerodynamic lift coefficient and the angle of attack under different flap angles is analyzed.
[0135] Step 2: Flap Signal Selection
[0136] In practice, the initial flap angle is set to 25°, the flap fault signal is set to 0 (1 for fault, 0 for normal), the total simulation time is 15s, and the flap angle changes from 25° to 17° at 5s.
[0137] Step 3: Generate compensation instructions
[0138] 301. In specific implementation, based on aerodynamic coefficient analysis, an interpolation table of angle of attack increments under different flaps is obtained, from... Figure 3 The interpolation table is as follows;
[0139] Flaps (°) Angle of attack increment (°) 0 0 17 -5 25 -8 40 -11
[0140] 302. In specific implementation, the interpolated angle of attack increment signal is passed through a high-pass filter to obtain the real-time pitch rate compensation command;
[0141] 303. In specific implementation, the pitch rate compensation command is limited and then multiplied by gain A to obtain the flap configuration change pitch rate compensation command;
[0142] 304. In specific implementation, the pitch rate compensation command is limited and then multiplied by the gain B to obtain the pitch rate increment command for flap configuration change;
[0143] Step 4: Instruction Synthesis
[0144] The flap configuration change pitch rate compensation command and the flap configuration change pitch rate increment command are respectively integrated into the longitudinal damping control loop and integral control loop of the main flight control system, thereby realizing compensation control during flap configuration changes. Figure 5 The figure shown is a schematic diagram of the simulation results of flap angle change compensation control in Embodiment 2 of the present invention.
[0145] pass Figure 4 and Figure 5 It can be seen that the flap configuration change compensation control method for large amphibious aircraft provided by the embodiments of the present invention can offset the adverse effects of configuration changes by compensating for pitch rate changes when flap angles change, thereby reducing the pilot's operational burden and improving safety to a certain extent.
[0146] While the embodiments disclosed in this invention are as described above, they are merely illustrative of the embodiments to facilitate understanding of the invention and are not intended to limit the invention. Any person skilled in the art to which this invention pertains may make any modifications and variations in the form and details of the implementation without departing from the spirit and scope disclosed herein; however, the scope of patent protection for this invention shall still be determined by the scope defined in the appended claims.
Claims
1. A method for compensating for flap configuration changes in a large amphibious aircraft, characterized in that, include: Step 1: Obtain the relationship between the aerodynamic lift coefficient and the angle of attack for the aircraft at different flap angles; Step 2: Select the appropriate flap signal based on whether the flap is malfunctioning; Step 3: Using the selected flap signal as the input variable, interpolation is performed based on the relationship between the aerodynamic lift coefficient and the angle of attack to obtain the angle of attack increment compensation value. By performing high-pass filtering and amplitude limiting on the angle of attack increment compensation value and multiplying it by different gains, the compensation commands for different longitudinal branches when the flap angle changes are obtained. Step 4: Connect the compensation commands for different longitudinal branches when the flap angle changes to the corresponding longitudinal control law to achieve the compensation function when the flap angle changes.
2. The flap configuration change compensation control method for large amphibious aircraft according to claim 1, characterized in that, Step 1 includes: Based on the aircraft configuration and lift coefficient curves under different flap angles, the relationship between the aerodynamic lift coefficient and the angle of attack under different flap angles is analyzed.
3. The flap configuration change compensation control method for large amphibious aircraft according to claim 1, characterized in that, Step 2 includes: The value of the flap signal is determined based on whether the flap is malfunctioning; When the flap signal is normal, select the current flap signal; when the flap signal is faulty, select the previous flap signal.
4. The flap configuration change compensation control method for large amphibious aircraft according to claim 1, characterized in that, The compensation commands for different longitudinal branches when the flap angle changes in step 3 include: the flap angle change pitch rate compensation command and the flap angle change pitch rate increment command.
5. The flap configuration change compensation control method for large amphibious aircraft according to any one of claims 1 to 4, characterized in that, Step 3 includes: Step 31: Using the selected flap signal as the input variable, obtain the change in angle of attack during different flap angle changes, and form an interpolation table of constant lift coefficient angle of attack increment. Step 32: Use the constant lift coefficient angle of attack increment interpolation table to obtain the angle of attack increment compensation value under different flap angles; Step 33: Pass the angle of attack increment compensation value command through a high-pass filter to obtain the pitch rate compensation command required for the angle of attack increment when the flap angle changes during the flap retraction and extension process. Step 34: After passing the pitch rate compensation command through the limiting circuit and multiplying it by the gain A, the pitch rate compensation command for flap angle change is obtained. Step 35: After passing the pitch rate compensation command through the limiting stage and multiplying it by the gain B, the pitch rate increment command for flap angle change is obtained.
6. The flap configuration change compensation control method for large amphibious aircraft according to claim 5, characterized in that, Step 31 involves forming an interpolation table for the constant lift coefficient angle of attack increment, including: Based on the aircraft lift coefficient curve, within the linear region of the basic lift coefficient, a constant lift coefficient straight line is determined for different flap angles. Based on the constant lift coefficient straight line, the change in angle of attack during the change of different flap angles is confirmed, forming a constant lift coefficient angle of attack increment interpolation table.
7. The flap configuration change compensation control method for large amphibious aircraft according to claim 6, characterized in that, The specific method for forming the constant lift coefficient angle-of-attack increment interpolation table in step 31 is as follows: Step 31a: For the lift coefficient curve, within the linear region of the basic lift coefficient, select the 0-degree flap lift coefficient curve corresponding to the specified angle of attack and draw a straight line with equal lift coefficient. Step 31b: Obtain the angle of attack value for each flap angle under the given constant lift coefficient straight line; Step 31c: Using the 0-degree flap as a reference, obtain the change in angle of attack corresponding to the transition from 0 degrees flap to each flap angle, which is the angle of attack increment compensation value under different flap angles. The specified angle of attack is selected within the range of angles of attack during a normal flap retraction / extension scenario.
8. The flap configuration change compensation control method for large amphibious aircraft according to claim 5, characterized in that, Step 4 includes: The flap angle change pitch rate increment command is superimposed with the aircraft pitch rate signal and then enters the integral control loop. The flap configuration change pitch rate compensation command is superimposed on the aircraft pitch rate signal and then enters the damping control loop.
9. A flap configuration change compensation control system for a large amphibious aircraft, characterized in that, The flap angle change compensation control system is used to execute the flap configuration change compensation control method for a large amphibious aircraft as described in any one of claims 1 to 8. The control includes: a lift coefficient relationship generation module, a flap signal selection module, a compensation command generation module, a compensation command output module, as well as an integral control loop and a damping control loop. The lift coefficient relationship generation module is used to obtain the relationship between the aerodynamic lift coefficient and the angle of attack of an aircraft at different flap angles; The flap signal selection module is used to select the appropriate flap signal based on whether the flap is malfunctioning. The compensation command generation module takes the flap signal selected by the flap signal selection module as the input variable, and interpolates the aerodynamic lift coefficient with the angle of attack obtained by the lift coefficient relationship generation module to obtain the angle of attack increment compensation value. By performing high-pass filtering and amplitude limiting on the angle of attack increment compensation value and multiplying it by different gains, the compensation commands for different longitudinal branches when the flap angle changes are obtained. The compensation commands for different longitudinal branches when the flap angle changes include: the flap angle change pitch rate compensation command and the flap angle change pitch rate increment command. The compensation command output module is used to input the compensation commands of different longitudinal branches when the flap angle changes into the corresponding longitudinal control law to realize the compensation function when the flap angle changes. Specifically, the flap angle change pitch rate increment command is superimposed with the aircraft pitch rate signal and then enters the integral control loop; the flap configuration change pitch rate compensation command is superimposed with the aircraft pitch rate signal and then enters the damping control loop.
10. A flap configuration change compensation control system for a large amphibious aircraft according to claim 9, characterized in that, The compensation instruction generation module specifically includes: The interpolation table unit is used to obtain the change in angle of attack during different flap angle changes by using the selected flap signal as the input variable, and to form a constant lift coefficient angle of attack increment interpolation table; it is also used to obtain the angle of attack increment compensation value under different flap angles using the constant lift coefficient angle of attack increment interpolation table. The filtering unit is used to pass the angle of attack increment compensation value command through a high-pass filter to obtain the pitch rate compensation command required for the angle of attack increment when the flap angle changes during the flap retraction and extension process. The limiting unit is used to pass the pitch rate compensation command through the limiting circuit and multiply it by gain A to obtain the flap angle change pitch rate compensation command; it is also used to pass the pitch rate compensation command through the limiting circuit and multiply it by gain B to obtain the flap angle change pitch rate increment command.