An aircraft high-altitude attitude stabilizing control method for air vane

Abstract

The application discloses an aircraft high-altitude attitude stabilizing control method for air rudders and relates to the technical field of aircraft attitude control. The method comprises the following steps: determining a three-loop feedback driver and a notch filter for the air rudders, wherein the three-loop feedback driver comprises a three-loop overload driver and a three-loop attitude driver; switching the three-loop overload driver into the three-loop attitude driver; reducing the outer loop gain parameter of the three-loop feedback driver to meet the requirements of the outer loop bandwidth and the stability margin; changing the input of the three-loop attitude driver in the pitch channel into the sum of the ballistic inclination angle and the preset attack angle of the aircraft, and changing the feedback into the pitch angle of the aircraft; and changing the input of the three-loop attitude driver in the yaw channel into the sum of the ballistic deflection angle and the sideslip angle of the aircraft, and changing the feedback into the yaw angle of the aircraft. The application realizes the stable control of the high-altitude attitude of the aircraft and has certain anti-interference performance.

Description

Technical Field

[0001] This invention relates to the field of aircraft attitude control technology, and specifically to an air-rudder-oriented high-altitude attitude stabilization control method for aircraft. Background Technology

[0002] With the rapid development of high technology, countries around the world are focusing on the development of aircraft technology. As aircraft technology continues to advance, penetration capabilities and precision strike capabilities have become key technologies in the field. Due to the increasingly complex and variable operating conditions of aircraft, advanced guidance technology is constantly evolving. Therefore, the requirements for aircraft attitude stabilization and control technology are becoming increasingly stringent.

[0003] Currently, most types of aircraft use aerodynamic control as the actuator to achieve stable attitude control. The problem with this method is that it is difficult to maintain stable flight attitude during high-altitude flight. Summary of the Invention

[0004] In view of the deficiencies in the existing technology, the technical problem solved by the present invention is: how to stabilize and control the flight attitude of an aircraft during high-altitude flight.

[0005] To achieve the above objectives, the present invention provides an air-rudder-oriented high-altitude attitude stability control method for aircraft. This method includes the following steps: determining an air-rudder-oriented three-loop feedback autopilot and a notch filter; the three-loop feedback autopilot includes a three-loop overload autopilot and a three-loop attitude autopilot; switching the three-loop overload autopilot to the three-loop attitude autopilot; reducing the outer loop gain parameter of the three-loop feedback autopilot to meet the requirements of outer loop bandwidth and stability margin; changing the input of the pitch channel three-loop attitude autopilot to the sum of the aircraft's trajectory tilt angle and a preset angle of attack, and changing the feedback to the aircraft's pitch angle; changing the input of the yaw channel three-loop attitude autopilot to the sum of the aircraft's trajectory deflection angle and sideslip angle, and changing the feedback to the aircraft's yaw angle.

[0006] Based on the above technical solution, the three-loop attitude control system's rudder control output in the pitch channel in, For pitch channel attitude angle deviation, ω z K is the pitch rate. wp K ip K np These are the inner loop gain, middle loop gain, and outer loop gain of the pitch channel three-loop attitude control system, respectively.

[0007] The three-circuit overload control system's rudder control output on the pitch channel in, For the trajectory tilt angle and angular rate deviation of the elevation channel, ω z K is the pitch rate. wp K ip K np These are the inner loop gain, middle loop gain, and outer loop gain of the three-loop overload control system for the pitch channel;

[0008] The rudder control output of the three-loop attitude control system in the yaw channel Where Δψ is the attitude angle deviation of the yaw channel, ω y For yaw rate, parameter K wy K iy K ny These represent the inner loop gain, middle loop gain, and outer loop gain of the three-loop attitude control system in the yaw channel, respectively.

[0009] The three-circuit overload control system's rudder control output in the yaw channel in, ω represents the ballistic deflection angle and angular rate deviation in the yaw path. y For yaw rate, parameter K wy K iy K ny These represent the inner loop gain, middle loop gain, and outer loop gain of the three-loop overload controller for the yaw channel, respectively.

[0010] Parameter K wp K ip K np The determination process is as follows: based on the dynamic coefficient design parameter K wp K ip K np The designed parameter K wp K ip K np Substitute the data into a mathematical simulation model for simulation verification;

[0011] Parameter K wy K iy K ny The determination process and parameter K wp K ip K np The process of determining them is the same.

[0012] Based on the above technical solution, the pitch rate ω z yaw rate ω y Determination: The pitch rate ω is obtained based on the elastic mathematical model. z yaw rate ω y .

[0013] Based on the above technical solution, the determination of the elastic mathematical model is as follows: an elastic mathematical model is established based on the coefficients of the elastic motion equation, aerodynamic parameters, and dynamic coefficients.

[0014] Based on the above technical solution, the dynamic coefficient is determined by obtaining the dynamic coefficient of the aircraft based on the aerodynamic derivative.

[0015] Based on the above technical solution, the aerodynamic derivative is determined by obtaining the aerodynamic derivative through aerodynamic parameters.

[0016] Based on the above technical solution, the determination of the elastic motion equation coefficients and the aerodynamic parameters involves: obtaining overall data, which includes the elastic motion equation coefficients and aerodynamic parameters.

[0017] Based on the above technical solution, the notch filter expression is as follows:

[0018]

[0019] Among them, input U in For the steering output of the three-loop feedback autopilot C k D k Here are the parameters of the notch filter network; the determined parameter value C k D k There are three groups, in the order of C0, C1, C2, D1, D2: 0.8401, -1.3808, 0.8307, -1.3808, 0.6708; 0.8816, -1.4490, 0.8717, -1.4490, 0.7532; 0.9781, -1.6077, 0.9671, -1.6077, 0.9452.

[0020] Based on the above technical solution, the high-altitude attitude stabilization control method for aircraft with air rudders further includes the following steps: performing mathematical simulation verification of the high-altitude attitude stabilization control method of the aircraft, setting various pull conditions through the mathematical simulation model to verify the correctness of the high-altitude attitude control method of the aircraft.

[0021] Compared with the prior art, the advantages of the present invention are as follows:

[0022] This invention addresses the shortcomings of insufficient aerodynamic rudder effectiveness at high altitudes, such as inadequate overload capacity and difficulty in tracking guidance system outputs. It rationally selects control methods and designs corresponding three-loop feedback autopilot and notch filter for aerodynamic rudders, achieving high-altitude attitude stability of the aircraft with a certain degree of anti-interference capability, without affecting the output of subsequent rudder control commands, thus meeting the attitude stability requirements and exhibiting good stability quality. Attached Figure Description

[0023] Figure 1 This is a flowchart of a high-altitude attitude stabilization control method for an aircraft with air rudders, as described in an embodiment of the present invention. Detailed Implementation

[0024] The present invention will be further described in detail below with reference to the accompanying drawings and embodiments.

[0025] See Figure 1 As shown, the high-altitude attitude stabilization control method for an aircraft with air rudders in this embodiment of the invention includes the following steps:

[0026] S1: Obtain the aerodynamic derivative through aerodynamic parameters;

[0027] S2: Calculate the dynamic coefficient of the aircraft based on the aerodynamic derivative;

[0028] S3: Establish an elastic mathematical model based on the coefficients of the elastic motion equation, aerodynamic parameters, and dynamic coefficients;

[0029] S4: Based on the elastic mathematical model, design a three-loop feedback autopilot for the aerodynamic rudder;

[0030] S5: Design a notch filter to filter out elastic interference in the three-loop feedback autopilot facing the aerodynamic rudder;

[0031] S6: Design a high-altitude attitude stabilization control method for an aircraft based on a three-loop feedback autopilot and a notch filter oriented towards the air rudder.

[0032] S7: Mathematical simulation verification of the high-altitude attitude stabilization control method for aircraft.

[0033] Building upon this, before obtaining the aerodynamic derivative from the aerodynamic parameters in S1, the following steps are included: acquiring overall data. The overall data includes the coefficients of the elastic motion equation and the aerodynamic parameters. The aerodynamic parameters are used to solve for the aerodynamic derivative in S2, and the coefficients of the elastic motion equation and the aerodynamic parameters are used to establish the elastic mathematical model in S3.

[0034] Based on this, the calculation process of the aerodynamic derivative in S1 is as follows:

[0035] (1) Obtain aerodynamic parameters from the overall data. Aerodynamic parameters include: the derivative of lift coefficient with respect to angle of attack, the derivative of lift coefficient with respect to rudder deflection, the derivative of pitch moment coefficient with respect to angle of attack, the derivative of pitch moment coefficient with respect to rudder deflection, the derivative of side force coefficient with respect to sideslip angle, the derivative of side force coefficient with respect to rudder deflection, the derivative of yaw moment coefficient with respect to sideslip angle, and the derivative of yaw moment coefficient with respect to rudder deflection.

[0036] (2) Based on the above data disclosure, those skilled in the art can process the above data (e.g., calculate the difference) to obtain the aerodynamic derivative according to existing knowledge.

[0037] Based on this, the calculation process of the dynamic coefficient in S2 is as follows:

[0038] Based on the derivatives of the pitch moment coefficient with respect to the angle of attack and the pitch moment coefficient with respect to the rudder deflection angle, as well as the relevant ballistic parameters, the stability dynamic coefficient (which reflects the static stability of the aircraft and is an indicator of its stability) and controllability dynamic coefficient (which reflects the control efficiency of the aircraft's actuators and is an indicator of its maneuverability) of the aircraft are calculated.

[0039] Based on this, the method for establishing the elastic mathematical model in S3 is as follows:

[0040] Based on all the dynamic coefficients obtained in S2 and the elastic motion equation coefficients and aerodynamic parameters obtained in the steps prior to S1, an elastic mathematical model is established to calculate the pitch rate ω. z yaw rate ω y .

[0041] Based on this, the three-loop feedback autopilot in S4 for aerodynamic control includes a three-loop attitude autopilot and a three-loop overload autopilot. Their output expressions are as follows:

[0042] Pitch channel three-loop attitude control system rudder control output in, The pitch channel attitude angle deviation (specifically, the difference between the standard pitch angle and the actual pitch angle), ω z For pitch rate, parameter K wp K ip K np These represent the inner loop gain, middle loop gain, and outer loop gain of the pitch channel three-loop attitude autopilot, respectively.

[0043] Pitch channel three-circuit overload control system rudder control output in ω represents the deviation of the ballistic tilt angle angular rate in the elevation channel (specifically, the difference between the standard ballistic tilt angle angular rate and the actual ballistic tilt angle angular rate). z For pitch rate, parameter K wp K ip K np These represent the inner loop gain, middle loop gain, and outer loop gain of the pitch channel three-loop overload control system, respectively.

[0044] Yaw channel three-loop attitude control system rudder control output Where Δψ is the yaw channel attitude angle deviation (specifically, the difference between the standard yaw angle and the actual yaw angle), ω y For yaw rate, parameter K wy K iy K nyThese represent the inner loop gain, middle loop gain, and outer loop gain of the three-loop attitude control system in the yaw channel, respectively.

[0045] Yaw channel three-circuit overload control system rudder control output in, ω represents the angular rate deviation of the ballistic deflection in the yaw path (specifically, the difference between the standard ballistic deflection angular rate and the actual ballistic deflection angular rate). y For yaw rate, parameter K wy K iy K ny These represent the inner loop gain, middle loop gain, and outer loop gain of the three-loop overload controller for the yaw channel, respectively.

[0046] From the above expression, it can be seen that, according to Parameter K wp K ip K np K wy K iy K ny And ω already calculated in S3 z ω y The rudder control output of the three-loop feedback autopilot oriented towards the aerodynamic rudder can be calculated.

[0047] Based on this, the three-loop feedback autopilot parameter K in S4 is oriented towards the aerodynamic rudder. wp K ip K np The process of determining is as follows:

[0048] S1: Design parameter K based on dynamic coefficient wp K ip K np ;

[0049] S2: The designed parameters K wp K ip K np The data is then substituted into a mathematical simulation model for verification. In the mathematical simulation, curves showing the changes of variables such as attitude angle, attitude angular rate, angle of attack, and sideslip angle over time can be viewed. If the simulation results show that these variables are within the specified normal range, the design is correct, and the simulation verification ends. If the simulation results exceed the specified normal range, the design does not meet the requirements, and the process proceeds to S401.

[0050] Parameter K wy K iy K ny The determination process and parameter K wp K ip K np The process of determining them is the same.

[0051] Based on this, S5 designs a notch filter to filter out elastic interference in the three-loop feedback autopilot facing the aerodynamic rudder. The notch filter expression is:

[0052]

[0053] Among them, input U in The steering output of the three-loop feedback autopilot calculated in S4 Output U out To filter out elastic interference, the rudder control output (the rudder control output U) out The parameter C is transmitted to the air rudder to control the aircraft's flight attitude. k D k The parameters for the filter network need to be selected based on frequency characteristics.

[0054] The parameters determined in this invention are in three groups, in the order of C0, C1, C2, D1, D2: 0.8401, -1.3808, 0.8307, -1.3808, 0.6708; 0.8816, -1.4490, 0.8717, -1.4490, 0.7532; 0.9781, -1.6077, 0.9671, -1.6077, 0.9452.

[0055] Based on this, S6 designs a high-altitude attitude stabilization control method for the aircraft using a three-loop feedback autopilot with air rudders and a notch filter. This mainly includes three steps: switching the autopilot, changing the outer loop gain parameters, and improving the three-loop attitude autopilot.

[0056] (1) Switching the driver's instrument:

[0057] At high altitudes, due to the low air density and insufficient dynamic pressure, the air rudder is not enough to generate the corresponding overload to meet the overload requirements of the guidance system. Therefore, it is necessary to switch the three-loop overload control system to a three-loop attitude control system.

[0058] (2) Change the outer loop gain parameter:

[0059] Due to the switching of the autopilot, the outer loop gain parameter needs to be changed accordingly. Specifically, the pitch channel outer loop gain parameter K needs to be adjusted. np Adjust the value to meet the requirements of outer loop bandwidth and stability margin; adjust the outer loop gain parameter K of the yaw channel. ny Adjust it to meet the requirements of outer ring bandwidth and stability margin.

[0060] (3) Improved three-loop attitude control system:

[0061] The inventors' improvement to the three-loop attitude control system is as follows: the input to the pitch channel three-loop attitude control system is the standard pitch angle calculated by the guidance system, and the feedback is the actual pitch angle and pitch rate of the aircraft; the input to the pitch channel three-loop overload control system is the standard ballistic tilt rate calculated by the guidance system, and the feedback is the actual ballistic tilt rate and pitch rate of the aircraft. Similarly, the input to the yaw channel three-loop attitude control system is the standard yaw angle calculated by the guidance system, and the feedback is the actual yaw angle and yaw rate of the aircraft; the input to the yaw channel three-loop overload control system is the standard ballistic deflection rate calculated by the guidance system, and the feedback is the actual ballistic deflection rate and yaw rate of the aircraft.

[0062] Since there may be some interference at high altitudes, it is relatively advantageous for an aircraft to fly with greater static stability at high altitudes for attitude stability. Based on aerodynamic parameters, the angle of attack (called the preset angle of attack) when the aircraft in the pitch channel is at maximum static stability can be calculated in advance, as can the sideslip angle (called the preset sideslip angle) when the aircraft in the yaw channel is at maximum static stability. Furthermore, high altitudes are generally in the maneuvering phase, making it difficult for the guidance system to calculate the standard pitch and yaw angles. Based on the three-loop attitude control system expression in S4, it is difficult to calculate the rudder control output of the three-loop attitude control system. It is difficult to maintain stable control over the high-altitude attitude of the aircraft.

[0063] The improvement to the three-loop attitude control system involves changing the input of the pitch channel's three-loop attitude control system to the sum of the aircraft's trajectory inclination angle and angle of attack, and changing the feedback to the aircraft's pitch angle. In this case, the angle for high-altitude control is the aircraft's angle of attack. When the aircraft's angle of attack reaches the preset value, the aircraft will fly through the high-altitude section with maximum static stability. Similarly, the input of the yaw channel's three-loop attitude control system is changed to the sum of the aircraft's trajectory deflection angle and the preset sideslip angle, and the feedback to the aircraft's yaw angle. In this case, the angle for high-altitude control is the aircraft's sideslip angle. When the aircraft's sideslip angle reaches the preset value, the aircraft will fly through the high-altitude section with maximum static stability. This satisfies the attitude control requirements while enhancing high-altitude reliability.

[0064] Based on this, the process of S7 conducting mathematical simulation verification of the high-altitude attitude stability control method for aircraft is as follows:

[0065] The correctness of the high-altitude attitude control method of the aircraft is verified by mathematical simulation model. Various pull conditions are set in the simulation test to verify the stability quality of the high-altitude attitude control method of the aircraft.

[0066] Therefore, this invention addresses the shortcomings of insufficient aerodynamic rudder effectiveness at high altitudes, such as inadequate overload capacity and difficulty in tracking guidance system outputs. It rationally selects control methods and designs corresponding three-loop feedback autopilot and notch filter for aerodynamic rudders, achieving high-altitude attitude stability of the aircraft with a certain degree of anti-interference capability. It does not affect the output of subsequent rudder control commands, meets the requirements for attitude stability, and has good stability quality.

[0067] Furthermore, this invention is not limited to the above-described embodiments. Those skilled in the art can make various improvements and modifications without departing from the principles of this invention, and these improvements and modifications are also considered within the scope of protection of this invention. Contents not described in detail in this specification are prior art known to those skilled in the art.

Claims

1. A high-altitude attitude stabilization control method for an aircraft oriented towards aerodynamic control, characterized in that, The method includes the following steps: determining a three-loop feedback autopilot and a notch filter oriented towards the aerodynamic control system, wherein the three-loop feedback autopilot includes a three-loop overload autopilot and a three-loop attitude autopilot; switching the three-loop overload autopilot to a three-loop attitude autopilot; reducing the outer loop gain parameter of the three-loop feedback autopilot to meet the requirements of outer loop bandwidth and stability margin; changing the input of the pitch channel three-loop attitude autopilot to the sum of the aircraft's trajectory tilt angle and preset angle of attack, and changing the feedback to the aircraft's pitch angle; changing the input of the yaw channel three-loop attitude autopilot to the sum of the aircraft's trajectory deflection angle and sideslip angle, and changing the feedback to the aircraft's yaw angle.

2. The high-altitude attitude stabilization control method for aircraft with air-rudder as described in claim 1, characterized in that, The three-loop attitude control system's rudder control output in the pitch channel in, For pitch channel attitude angle deviation, ω z K is the pitch rate. wp K ip K np These are the inner loop gain, middle loop gain, and outer loop gain of the pitch channel three-loop attitude control system, respectively. The three-circuit overload control system's rudder control output on the pitch channel in, For the trajectory tilt angle and angular rate deviation of the elevation channel, ω z K is the pitch rate. wp K ip K np These are the inner loop gain, middle loop gain, and outer loop gain of the three-loop overload control system for the pitch channel; The rudder control output of the three-loop attitude control system in the yaw channel Where Δψ is the attitude angle deviation of the yaw channel, ω y For yaw rate, parameter K wy K iy K ny These represent the inner loop gain, middle loop gain, and outer loop gain of the three-loop attitude control system in the yaw channel, respectively. The three-circuit overload control system's rudder control output in the yaw channel in, ω represents the ballistic deflection angle and angular rate deviation in the yaw path. y For yaw rate, parameter K wy K iy K ny These represent the inner loop gain, middle loop gain, and outer loop gain of the three-loop overload controller for the yaw channel, respectively. Parameter K wp K ip K np The determination process is as follows: based on the dynamic coefficient design parameter K wp K ip K np The designed parameter K wp K ip K np Substitute the data into a mathematical simulation model for simulation verification; Parameter K wy K iy K ny The determination process and parameter K wp K ip K np The process of determining them is the same.

3. The high-altitude attitude stabilization control method for aircraft with air rudders as described in claim 2, characterized in that, The pitch rate ω z yaw rate ω y Determination: The pitch rate ω is obtained based on the elastic mathematical model. z yaw rate ω y .

4. The high-altitude attitude stabilization control method for aircraft with air rudders as described in claim 3, characterized in that, The determination of the elastic mathematical model: The elastic mathematical model is established based on the coefficients of the elastic motion equation, aerodynamic parameters, and dynamic coefficients.

5. The high-altitude attitude stabilization control method for aircraft with air-rudder as described in claim 4, characterized in that, The dynamic coefficient is determined by obtaining the dynamic coefficient of the aircraft based on the aerodynamic derivative.

6. The high-altitude attitude stabilization control method for an aircraft with air-rudder as described in claim 5, characterized in that, The aerodynamic derivative is determined by obtaining the aerodynamic parameters.

7. The high-altitude attitude stabilization control method for an aircraft with air-rudder as described in claim 6, characterized in that, Determination of the elastic motion equation coefficients and the aerodynamic parameters: Obtain overall data, which includes the elastic motion equation coefficients and aerodynamic parameters.

8. The high-altitude attitude stabilization control method for an aircraft with air-rudder as described in claim 7, characterized in that, The notch filter expression is as follows: Among them, input U in For the steering output of the three-loop feedback autopilot C k D k Here are the parameters of the notch filter network; the determined parameter value C k D k There are three groups, in the order of C0, C1, C2, D1, D2: 0.8401, -1.3808, 0.8307, -1.3808, 0.6708; 0.8816, -1.4490, 0.8717, -1.4490, 0.7532; 0.9781, -1.6077, 0.9671, -1.6077, 0.9452.

9. The high-altitude attitude stabilization control method for an aircraft with air rudders as described in claim 8, characterized in that, The method also includes the following steps: mathematical simulation verification of the high-altitude attitude stabilization control method of the aircraft, and setting various pull conditions through the mathematical simulation model to verify the correctness of the high-altitude attitude control method of the aircraft.

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